Operator's Manual For UH 60A Helicopter TM 1 1520 237 10

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*TM 1-1520-237-10
TECHNICAL MANUAL

OPERATOR’S MANUAL
FOR
UH-60A HELICOPTER
UH-60L HELICOPTER
EH-60A HELICOPTER

DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited.
*This manual supersedes TM 1-1520-237-10, dated 31 August 1994, including all changes.

HEADQUARTERS, DEPARTMENT OF THE ARMY
31 OCTOBER 1996

TM 1--1520--237--10
10

CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 September 2002

NO. 10

OPERATOR’S MANUAL
FOR

UH-60A, UH-60L AND EH-60A HELICOPTERS
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 1--1520--237--10, dated 31 October 1996, is changed as follows:
1. Title has been changed to read as stated above.
2. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar
in the margin. An illustration change is indicated by the current change number. Text that flows to the
following page is indicated by the current change number.

Remove pages
A through D
i through v/(vi Blank)
2--1 and 2--2
2--11 and 2--12
2--15 and 2--16
2--21 through 2--24
2--31 through 2--34
2--34.1/(2--34.2 Blank)
2--35 and 2--36
2--45 through 2--48
-----------------------------2--55 and 2--56
2--64.1 and 2--64.2
2--83 and 2--84
2--93 and 2--94
3--7 and 3--8
3--11 and 3--12
3--19 through 3--26
-----------------------------3--33 and 3--34

Insert pages
A through D
i through v/(vi Blank)
2--1 and 2--2
2--11 and 2--12
2--15 and 2--16
2--21 through 2--24
2--31 through 2--34
2--34.1/(2--34.2 Blank)
2--35 and 2--36
2--45 through 2--48
2--48.1/(2--48.2 Blank)
2--55 and 2--56
2--64.1 and 2--64.2
2--83 and 2--84
2--93 and 2--94
3--7 and 3--8
3--11 and 3--12
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3--26.1 through 3--26.3/(3--26.4 Blank)
3--33 and 3--34

TM 1--1520--237--10
C10

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3--46.3 through 3--46.8
3--46.11 through 3--46.18
3--55 and 3--56
3--73 through 3--76
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4--12.1/(4--12.2 Blank)
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4--35 through 4--40
4--42.1/(4--42.2 Blank)
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7A--143 and 7A--144
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8--10.1/(8--10.2 Blank)
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8--16.1/(8--16.2 Blank)
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-----------------------------9--13 through 9--18
9--18.1/(9--18.2 Blank)
9--19 through 9--24

Insert pages
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3--46.11 through 3--46.18
3--55 and 3--56
3--73 through 3--76
4--3 through 4--11/(4--12 Blank)
-----------------------------4--25 and 4--26
4--35 through 4--40
4--42.1/(4--42.2 Blank)
4--48.1/(4--48.2 Blank)
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4--50.1/(4--50.2 Blank)
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4--59 and 4--60
4--60.1 through 4--60.6
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5--23 through 5--25/(5--26 Blank)
6--7 through 6--12
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7A--17 and 7A--18
7A--143 and 7A--144
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8--10.1/(8--10.2 Blank)
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8--16.1/(8--16.2 Blank)
8--17 through 8--22
8--22.1/(8--22.2 Blank)
9--13 through 9--18
9--18.1/(9--18.2 Blank)
9--19 through 9--24

3. Retain these sheets in front of manual for reference purposes.

TM 1--1520--237--10
10

By Order of the Secretary of the Army:

ERIC K. SHINSEKI
General, United States Army
Chief of Staff
Official:

JOEL B. HUDSON
Administrative Asssistant to the
Secretary of the Army
0217923

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284, requirements for
TM 1--1520--237--10.

TM 1--1520--237--10
C9

CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 19 April 2002

NO. 9

OPERATOR’S MANUAL
FOR

UH-60A, UH-60L AND EH-60A HELICOPTERS
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 1--1520--237--10, dated 31 October 1996, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar
in the margin. An illustration change is indicated by the current change number. Text that flows to the
following page is indicated by the current change number.

Remove pages
A through D
i through iv
2--1 and 2--2
2--11 and 2--12
2--25 through 2--28
2--31 and 2--32
2--41 through 2--44
2--57 and 2--58
2--61 and 2--62
2--64.1 and 2--64.2
2--67 and 2--68
2--77 and 2--78
2--81 through 2--86
2--89 and 2--90
2--93 through 2--96
3--8.1 and 3--8.2
3--32.7 and 3--32.8
3--35 and 3--36
3--46.7 through 3--46.12
3--65 and 3--66
4--31 and 4--32

Insert pages
A through D
i through iv
v/(vi Blank)
2--1 and 2--2
2--11 and 2--12
2--25 through 2--28
2--31 and 2--32
2--41 through 2--44
2--57 and 2--58
2--61 and 2--62
2--64.1 and 2--64.2
2--67 and 2--68
2--77 and 2--78
2--81 through 2--86
2--89 and 2--90
2--93 through 2--96
3--8.1 and 3--8.2
3--32.7 and 3--32.8
3--35 and 3--36
3--46.7 through 3--46.12
3--65 and 3--66
4--31 and 4--32

TM 1--1520--237--10
C9

Remove pages
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4--49 and 4--50
4--50.1/(4--50.2 Blank)
4--55 and 4--56
4--67/(4--68 Blank)
5--5 through 5--8
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5--21 through 5--25/(5--26 Blank)
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-------------------------------8--13 through 8--16
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9--15 and 9--16

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4--49 and 4--50
4--50.1/(4--50.2 Blank)
4--55 and 4--56
4--67/(4--68 Blank)
5--5 through 5--8
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5--21 through 5--25/(5--26 Blank)
8--3 and 8--4
8--5 and 8--6
8--9 and 8--10
8--10.1/(8.10.2 Blank)
8--13 through 8--16
9--11 and 9--12
9--15 and 9--16

2. Retain this sheet in front of manual for reference purposes.

By Order of the Secretary of the Army:

ERIC K. SHINSEKI
General, United States Army
Chief of Staff
Official:

JOEL B. HUDSON
Administrative Asssistant to the
Secretary of the Army
0122803

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284, requirements for
TM 1--1520--237--10.

TM 1-1520-237-10
C8
CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 15 JUNE 2001

NO. 8

OPERATOR’S MANUAL
FOR
UH-60A HELICOPTERS, UH-60L HELICOPTERS
AND EH-60A HELICOPTERS
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 1-1520-237-10, dated 31 October 1996, is changed as follows:
1.

Remove and insert pages as indicated below. New or changed text material is indicated by a vertical
bar in the margin. An illustration change is indicated by a vertical bar next to the figure title. Text that
flows to the following page is indicated by a current change number.

Remove pages
a and b
A through C/(D Blank)
i through iv
2-1 and 2-2
2-11 and 2-12
2-12.1/(2-12.2 Blank)
2-15 and 2-16
2-25 and 2-26
2-29 and 2-30
2-33 and 2-34
2-34.1/(2-34.2 Blank)
2-35 through 2-40
2-53 and 2-54
2-55 through 2-58
2-85 and 2-86
3-1 through 3-4
3-7 and 3-8
3-8.1 and 3-8.2
3-29 and 3-30

Insert pages
a and b
A through D
i through iv
2-1 and 2-2
2-11 and 2-12
2-12.1/(2-12.2 Blank)
2-15 and 2-16
2-25 and 2-26
2-29 and 2-30
2-33 and 2-34
2-34.1/(2-34.2 Blank)
2-35 through 2-40
2-53 and 2-54
2-55 through 2-58
2-85 and 2-86
3-1 through 3-4
3-7 and 3-8
3-8.1 and 3-8.2
3-29 and 3-30

TM 1-1520-237-10
C8

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3-32.1 through 3-32.11/(3-32.12 Blank)
3-35 and 3-36
3-46.3 and 3-46.4
3-46.7 through 3-46.10
3-46.21 and 3-46.22
3-65 and 3-66
3-73 and 3-74
3-77 and 3-78
3-78.1 and 3-78.2
3-79/(3-80 Blank)
4-9 through 4-12
4-25 and 4-26
-------------4-51 through 4-54
5-5 and 5-6
5-19 and 5-20
6-3 through 6-6
7-55 and 7-56
7-125 and 7-126
7-137 and 7-138
7A-139 and 7A-140
7A-143 and 7A-144
8-3 and 8-4
8-4.1/(8-4.2 Blank)
8-5 through 8-8
8-8.1/(8-8.2 Blank)
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8-15 and 8-16
8-17 through 8-22
8-23/(8-24 Blank)
9-1 through 9-6
9-6.1/(9-6.2 Blank)
9-11 through 9-18
9-19 and 9-20
B-1 through B-3/(B-4 Blank)
--------------

Insert pages
3-32.1 through 3-32.13/(3-32.14 Blank)
3-35 and 3-36
3-46.3 and 3-46.4
3-46.7 through 3-46.10
3-46.21 and 3-46.22
3-65 and 3-66
3-73 and 3-74
3-77 and 3-78
3-78.1 and 3-78.2
3-79/(3-80 Blank)
4-9 through 4-12
4-25 and 4-26
4-26.1/(4-26.2 Blank)
4-51 through 4-54
5-5 and 5-6
5-19 and 5-20
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7-55 and 7-56
7-125 and 7-126
7-137 and 7-138
7A-139 and 7A-140
7A-143 and 7A-144
8-3 and 8-4
8-4.1/(8-4.2 Blank)
8-5 through 8-8
8-8.1/(8-8.2 Blank)
8-9 through 8-12
8-15 and 8-16
8-17 through 8-22
8-23/(8-24 Blank)
9-1 through 9-6
9-6.1/(9-6.2 Blank)
9-11 through 9-18
9-19 and 9-20
B-1 through B-3/(B-4 Blank)
C-1 through C-3/(C-4 Blank)

TM 1-1520-237-10
C8

By Order of the Secretary of the Army:
Official:
ERIC K. SHINSEKI
General, United States Army
Chief of Staff

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
0103660

DISTRIBUTION:
To be distributed in accordance with initial distribution number (IDN) 310284 requirements for
TM 1-1520-237-10.

TM 1--1520--237--10
C7

CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 27 NOVEMBER 2000

NO. 7

OPERATOR’S MANUAL
FOR

UH--60A HELICOPTERS, UH--60L HELICOPTERS
AND EH--60A HELICOPTERS

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 1--1520--237--10, dated 31 October 1996, is changed as follows:
1.

Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar
in the margin. An illustration change is indicated by the current change number. Text that flows to the
following page is indicated by the current change number.

Remove pages

Insert pages

A through C/(D Blank)
2--1 and 2--2
2--21 and 2--22
2--47 and 2--48
2--87 and 2--88

2--93 and 2--94
4--5 and 4--6
4--9 through 4--12
4--43 and 4--44
4--47 and 4--48

4--48.1/(4--48.2 Blank)
4--51 through 4--54
4--60.5 and 4--60.6
4--65 and 4--66
7A--15 and 7A--16

9--1 and 9--2

TM 1-1520-237-10
C7

Remove pages

Insert pages

9-6.1/(9-6.2 Blank)
9-19 through 9-22
B-1 and B-2

2. Retain this sheet in front of manual for reference purposes.

By Order of the Secretary of the Army:
Official:
ERIC K. SHINSEKI
General, United States Army
Chief of Staff

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
0019571

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284 requirements for
TM 1-1520-237-10.

TM 1--1520--237--10
C6

CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 3 APRIL 2000

NO. 6

OPERATOR’S MANUAL
FOR

UH--60A HELICOPTERS, UH--60L HELICOPTERS
AND EH--60A HELICOPTERS

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 1--1520--237--10, dated 31 October 1996, is changed as follows:
1.

Remove pages

Insert pages

A through C/(D Blank)
1--1 and 1--2
2--11 and 2--12
2--12.1/(2--12.2 Blank)
2--33 and 2--34

2--34.1/(2--34.2 Blank)
2--47 and 2--48
2--51 through 2--54
2--77 and 2--78
2--89 and 2--90

2--34.1/(2--34.2 Blank)
2--47 and 2--48
2--51 through 2--54
2--77 and 2--78
2--89 and 2--90
4-9 through 4-12

4-9 through 4-12
4--41 and 4--42
-----------------------------------------4--53 and 4--54
4--67/(4--68 Blank)
5--1 and 5--2

4--41 and 4--42
4--42.1/(4--42.2 Blank)
4--53 and 4--54
4--67/(4--68 Blank)
5--1 and 5--2

TM 1-1520-237-10
C6

Remove pages
5-9 and 5-10
5-23 and 5-24
6-1 and 6-2
7-1 and 7-2
-------------------7A-5 and 7A-6
-------------------7A-7 through 7A-12
-------------------7A-15 and 7A-16
8-7 and 8-8
-------------------8-11 and 8-12
9-19 through 9-22
A-1 and A-2

Insert pages
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7-2.1/(7-2.2 Blank)
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7A-7 through 7A-12
7A-12.1/(7A-12.2 Blank)
7A-15 and 7A-16
8-7 and 8-8
8-8.1/(8-8.2 Blank)
8-11 and 8-12
9-19 through 9-22
A-1 and A-2

2. Retain these sheets in front of the manual for reference purposes.

By Order of the Secretary of the Army:
ERIC K. SHINSEKI
General, United States Army
Chief of Staff
OFFICIAL:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
9931322

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284,
requirements for TM 1-1520-237-10.

TM 1-1520-237-10
C5
CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 JULY 1999

NO. 5

OPERATOR’S MANUAL
FOR
UH-60A HELICOPTERS, UH-60L HELICOPTERS
AND EH-60A HELICOPTERS
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.

TM 1-1520-237-10, dated 31 October 1996, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a
vertical bar in the margin. An illustration change is indicated by the current change number.
Text that flows to the following page is indicated by the current change number.
Remove pages
None
2-11 and 2-12
2-25 and 2-26
2-35 and 2-36
2-45 and 2-46
2-51 through 2-54
2-54.1/(2-54.2 blank)
2-55 and 2-56
2-73 and 2-74
2-77 through 2-80
2-83 and 2-84
2-89 through 2-92
2-95 and 2-96
3-5 through 3-8
3-8.1/(3-8.2 blank)
3-15 and 3-16

Insert pages
A through C/(D blank)
2-11 and 2-12
2-25 and 2-26
2-35 and 2-36
2-45 and 2-46
2-51 through 2-54
2-54.1/(2-54.2 blank)
2-55 and 2-56
2-73 and 2-74
2-77 through 2-80
2-83 and 2-84
2-89 through 2-92
2-95 and 2-96
3-5 through 3-8
3-8.1/(3-8.2 blank)
3-15 and 3-16

TM 1-1520-237-10
C5
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3-53 and 3-54
None
3-69 through 3-72
None
3-75 and 3-76
4-11 and 4-12
4-48.1/(4-48.2 blank)
4-49 and 4-50
None
5-19 and 5-20
6-23 and 6-24
7A-3 and 7A-4
8-3 and 8-4
8-5 through 8-10
8-13 through 8-16
8-17 through 8-20
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9-19 through 9-22
Index-3 through Index-8
None
Index-13 and Index-14
None
Index-19 through Index-22
None

Insert pages
3-53 and 3-54
3-54.1/(3-54.2 blank)
3-69 through 3-72
3-72.1/(3-72.2 blank)
3-75 and 3-76
4-11 and 4-12
4-48.1/(4-48.2 blank)
4-49 and 4-50
4-50.1/(4-50.2 blank)
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7A-3 and 7A-4
8-3 and 8-4
8-5 through 8-10
8-13 through 8-16
8-17 through 8-20
9-17 and 9-18
9-19 through 9-22
Index-3 through Index-8
Index-8.1/(Index-8.2 blank)
Index-13 and Index-14
Index-14.1/(Index-14.2 blank)
Index-19 through Index-22
Index-22.1/(Index-22.2 blank)

2. Retain these sheets in front of the manual for reference purposes.
By Order of the Secretary of the Army:
ERIC K. SHINSEKI
General, United States Army
Chief of Staff
OFFICIAL:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
9917308

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 313721,
requirements for TM 1-1520-237-10.

TM 1-1520-237-10
C4
CHANGE
NO. 4

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 29 JANUARY 1999

OPERATOR’S MANUAL
FOR
UH-60A HELICOPTERS, UH-60L HELICOPTERS, AND EH-60A HELICOPTERS
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 1-1520-237-10, dated 31 October 1996, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a
vertical bar in the margin. An illustration change is indicated by a vertical bar next to the figure
title. Text that flows to the following page is indicated by the current change number
Remove pages

Insert pages

a/(b blank)
1-1 and 1-2
2-1 and 2-2
2-7 and 2-8
2-11 and 2-12
2-12.1/(2-12.2 blank)
2-17 through 2-20
2-63 and 2-64
3-3 through 3-8
3-29 through 3-32
None
3-35 and 3-36
3-45 and 3-46
3-46.3 and 3-46.4
3-46.7 through 3-46.12
3-46.15 through 3-46.20
3-65 and 3-66
3-69 and 3-70
3-78.1 and 3-78.2
4-11 and 4-12
None

a and b
1-1 and 1-2
2-1 and 2-2
2-7 and 2-8
2-11 and 2-12
2-12.1/(2-12.2 blank)
2-17 through 2-20
2-63 and 2-64
3-3 through 3-8
3-29 through 3-32
3-32.1 through 3-32.11/(3-32.12 blank)
3-35 and 3-36
3-45 and 3-46
3-46.3 and 3-46.4
3-46.7 through 3-46.12
3-46.15 through 3-46.20
3-65 and 3-66
3-69 and 3-70
3-78.1 and 3-78.2
4-11 and 4-12
4-12.1/(4-12.2 blank)

TM 1-1520-237-10
C4
Remove pages
4-15 and 4-16

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4-15 and 4-16

4-19 and 4-20
4-51 through 4-54
4-57 through 4-60
None
5-23 and 5-24
None
8-7 through 8-16
None
8-23/(8-24 blank)
9-23 and 9-24
None
A-1 and A-2
B-1 through B-3/B-4 blank)

4-19 and 4-20
4-51 through 4-54
4-57 through 4-60
4-60.1 through 4-60.6
5-23 and 5-24
5-25/(5-26 blank)
8-7 through 8-16
8-16.1/(8-16.2 blank)
8-23/(8-24 blank)
9-23 and 9-24
9-25/(9-26 blank)
A-1 and A-2
B-1 through B-3/(B-4 blank)

2. Retain these sheets in front of the manual for reference purposes.

By Order of the Secretary of the Army:
DENNIS J. REIMER
General, United States Army
Chief of Staff
OFFICIAL:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
05406

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284,
requirements for TM 1-1520-237-10.

TM 1-1520-237-10
C3
CHANGE
NO. 3

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 OCTOBER 1998

Insert pages

a/(b blank)
i and ii
2-33 and 2-34
None
2-43 and 2-44
2-47 and 2-48
2-63 and 2-64
2-69 and 2-70
2-73 through 2-76
2-91 and 2-92
4-53 through 4-56
4-65 and 4-66
5-1 and 5-2
5-5 through 5-8
5-23 and 5-24
None
6-15 and 6-16
7-3 through 7-6
8-3 and 8-4
None
8-5 through 8-16
8-19 through 8-22

a and b
i and ii
2-33 and 2-34
2-34.1/(2-34.2 blank)
2-43 and 2-44
2-47 and 2-48
2-63 and 2-64
2-69 and 2-70
2-73 through 2-76
2-91 and 2-92
4-53 through 4-56
4-65 and 4-66
5-1 and 5-2
5-5 through 5-8
5-23 and 5-24
5-25/(5-26 blank)
6-15 and 6-16
7-3 through 7-6
8-3 and 8-4
8-4.1/(8-4.2 blank)
8-5 through 8-16
8-19 through 8-22

TM 1-1520-237-10
C3
Remove pages
8-23/(8-24 blank)
9-6.1/(9-6.2 blank)
9-11 and 9-12
9-15 through 9-18
None
9-19 through 9-24
A-1 and A-2
B-1 and B-2
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B-1 and B-2
B-3/(B-4 blank)

2. Retain these sheets in front of the manual for reference purposes.

By Order of the Secretary of the Army:
DENNIS J. REIMER
General, United States Army
Chief of Staff
OFFICIAL:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
05178

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284,
requirements for TM 1-1520-237-10.

TM 1-1520-237-10
C2
CHANGE

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 29 MAY 1998

NO. 2

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TM 1-1520-237-10
C2
2. Retain these sheets in front of the manual for reference purposes.

By Order of the Secretary of the Army:
DENNIS J. REIMER
General, United States Army
Chief of Staff
OFFICIAL:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
04944

DISTRIBUTION:
To be distributed in accordance with Initial Distribution Number (IDN) 310284,
requirements for TM 1-1520-237-10.

TM 1-1520-237-10
C1
CHANGE
NO. 1

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 JUNE 1997

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TM 1-1520-237-10
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2. Retain these sheets in front of the manual for reference purposes.

By Order of the Secretary of the Army:
DENNIS J. REIMER
General, United States Army
Chief of Staff
OFFICIAL:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
03940

DISTRIBUTION:
To be distributed in accordance with DA Form 12-31 E, block no. 0284, requirements for
TM 1-1520-237-10.

*TM 1-1520-237-10

WARNING

Personnel performing operations, procedures, and practices which are included or implied in this technical manual shall
observe the following warnings. Disregard of these warnings and precautionary information can cause serious injury or loss
of life.
BATTERY ELECTROLYTE
Battery electrolyte is harmful to the skin and clothing. If potassium hydroxide is spilled on clothing or other material, wash
immediately with clean water. If spilled on personnel, immediately flush the affected area with clean water. Continue
washing until medical assistance arrives. Neutralize any spilled electrolyte by thoroughly flushing contacted area with water.
CARBON MONOXIDE
When smoke, suspected carbon monoxide fumes, or symptoms of anoxia exist, the crew should immediately ventilate the
cockpit.
ELECTROMAGNETIC INTERFERENCE (EMI)
No electrical/electronic devices of any sort, other than those described in this manual or appropriate airworthiness release and
approved by USAATCOM AMSAT-R-ECU, are to be operated by crewmembers or passengers during operation of this
helicopter.
FIRE EXTINGUISHER
Exposure to high concentrations of extinguishing agent or decomposition products should be avoided. The liquid should not
be allowed to come into contact with the skin, as it may cause frost bite or low temperature burns.
HANDLING FUEL AND OIL
Turbine fuels and lubricating oils contain additives which are poisonous and readily absorbed through the skin. Do not allow
them to remain on skin longer than necessary.
HIGH VOLTAGE
All ground handling personnel shall be informed of high voltage hazards when making external cargo hookups.
NOISE
Sound pressure levels in this helicopter during some operating conditions exceed the Surgeon General’s hearing conservation
criteria, as defined in DA PAM 40-501. Hearing protection devices, such as the aviator helmet or ear plugs are required to
be worn by all personnel in and around the helicopter during its operation. When window guns are firing, when flights exceed
100 minutes during any 24 hour period, or when speeds are above 120 knots, helmet and ear plugs shall be worn by all
crewmembers.
WEAPONS AND AMMUNITION
Observe all standard safety precautions governing the handling of weapons and live ammunition. When not in use, point all
weapons in a direction offering the least exposure to personnel and property in case of accidental firing. Do not walk in front
of weapons. SAFE the machinegun before servicing. To avoid potentially dangerous situations, follow all procedural
warnings in text.
ELECTROMAGNETIC RADIATION
Do not stand within six feet of Aircraft Survivability Equipment (ASE), ALQ-156, ALQ-162, and ALQ-144 transmit
antennas when the ASE equipment is on. High frequency electromagnetic radiation can cause internal burns without causing
any sensation of heat. The HF radio transmits high power electromagnetic radiation. Serious injury or death can occur if you
Change 8

*TM 1-1520-237-10

touch the HF antenna while it is transmitting. Do not grasp, or lean against the antenna when power is applied to the
helicopter.
ALQ-144
Do not continuously look at the ALQ-144 infrared countermeasure transmitter during operation, or for a period of over 1
minute from a distance of less than 3 feet. Skin exposure to countermeasure radiation for longer than 10 seconds at a distance
less than 4 inches shall be avoided.

Change 8

TM 1-1520-237-10

LIST OF EFFECTIVE PAGES

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NOTE: On a changed page, the portion of the text affected by the latest change is indicated by a vertical line
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1

TM 1-1520-237-10

TECHNICAL MANUAL

HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C. 31 OCTOBER 1996

NO. 1-1520-237-10

Operator’s Manual
for
UH-60A, UH-60L, EH-60A HELICOPTERS
REPORTING ERRORS AND RECOMMENDING IMPROVEMENTS
You can help improve this manual. If you find any mistakes, or if you know of a way to improve
these procedures, please let us know. Mail your letter, DA Form 2028 (Recommended Changes to
Publications and Blank Forms), or DA Form 2028-2 located in the back of this manual, direct to:
Commander, US Army Aviation and Missile Command, ATTN: AMSAM-MMC-MA-NP, Redstone
Arsenal, AL 35898-5000. A reply will be furnished to you. You may also provide DA Form 2028
information to AMCOM via e-mail, fax, or the World Wide Web. Our fax number is: DSN 788-6546
or Commercial 256-842-6546. Our e-mail address is: 2028@redstone.army.mil. Instructions for
sending an electronic 2028 may be found at the back of this manual immediately preceding the
hard copy 2028. For the World Wide Web use: https://amcom2028.redstone.army.mil.
DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited.

NOTE
This document has been reviewed for the presence of Class I Ozone Depleting Chemicals. As of
Change 4, dated 29 January 1999, all references to Class I Ozone Depleting Chemicals have been
removed from this document by substitution with chemicals that do not cause atmospheric ozone
depletion.
TABLE OF CONTENTS
Chapter
&
Section

Page

CHAPTER 1

INTRODUCTION .........................................................................................

1-1

CHAPTER 2

AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION .........

2-1

Section I

Aircraft...........................................................................................................

2-1

Section II

Emergency Equipment ..................................................................................

2-25

Section III

Engines and Related Systems .......................................................................

2-27

Section IV

Fuel System ..................................................................................................

2-39

Section V

Flight Controls...............................................................................................

2-41

Change 10

TM 1-1520-237-10

TABLE OF CONTENTS (Cont)
Chapter
&
Section
Section VI

Hydraulic and Pneumatic System .................................................................

2-48.1

Section VII

Powertrain System.........................................................................................

2-53

Section VIII

Main and Tail Rotor Groups.........................................................................

2-55

Section IX

Utility Systems ..............................................................................................

2-57

Section X

Heating, Ventilating, Cooling, and Environmental Control Unit ................

2-62

Section XI

Electrical Power Supply and Distribution Systems......................................

2-64

Section XII

Auxiliary Power Unit ....................................................................................

2-71

Section XIII

Lighting..........................................................................................................

2-74

Section XIV

Flight Instruments..........................................................................................

2-77

Section XV

Servicing, Parking, and Mooring ..................................................................

2-86

AVIONICS ....................................................................................................

3-1

Section I

General...........................................................................................................

3-1

Section II

Communications ............................................................................................

3-8

Section III

Navigation......................................................................................................

3-32.11

Section IV

Transponder and Radar .................................................................................

3-75

MISSION EQUIPMENT...............................................................................

4-1

Section I

Mission Avionics...........................................................................................

4-1

Section II

Armament ......................................................................................................

4-26

Section III

Cargo Handling Systems...............................................................................

4-42

Section IV

Mission Flexible Systems..............................................................................

4-47

OPERATING LIMITS AND RESTRICTIONS...........................................

5-1

CHAPTER 3

CHAPTER 4

CHAPTER 5

Page

Change 10

TM 1-1520-237-10

TABLE OF CONTENTS (Cont)
Chapter
&
Section

Page

Section I

General...........................................................................................................

5-1

Section II

System Limits ................................................................................................

5-2

Section III

Power Limits..................................................................................................

5-9

Section IV

Loading Limits ..............................................................................................

5-12

Section V

Airspeed Limits .............................................................................................

5-13

Section VI

Maneuvering Limits ......................................................................................

5-19

Section VII

Environmental Restrictions ...........................................................................

5-23

Section VIII

Other Limitations...........................................................................................

5-24

WEIGHT/BALANCE AND LOADING ......................................................

6-1

Section I

General...........................................................................................................

6-1

Section II

Weight and Balance ......................................................................................

6-3

Section III

Fuel/Oil ..........................................................................................................

6-5

Section IV

Personnel........................................................................................................

6-7

Section V

Mission Equipment........................................................................................

6-13

Section VI

Cargo Loading ...............................................................................................

6-19

Section VII

Center of Gravity...........................................................................................

6-24

PERFORMANCE DATA..............................................................................

7-1

Section I

Introduction....................................................................................................

7-1

Section II

Maximum Torque Available .........................................................................

7-6

Section III

Hover..............................................................................................................

7-9

Section IV

Cruise .............................................................................................................

7-13

Section V

Optimum Cruise ............................................................................................

7-132

CHAPTER 6

CHAPTER 7

Change 10

iii

TM 1-1520-237-10

TABLE OF CONTENTS (Cont)
Chapter
&
Section

Section VI

Drag................................................................................................................

7-135

Section VII

Climb - Descent.............................................................................................

7-138

Section VIII

Fuel Flow.......................................................................................................

7-141

Section IX

Airspeed System Characteristics...................................................................

7-143

Section X

Special Mission Performance........................................................................

7-147

701C .....................................................................

7A-1

Section I

Introduction....................................................................................................

7A-1

Section II

Maximum Torque Available .........................................................................

7A-6

Section III

Hover..............................................................................................................

7A-13

Section IV

Cruise .............................................................................................................

7A-17

Section V

Optimum Cruise ............................................................................................

7A-134

Section VI

Drag................................................................................................................

7A-137

Section VII

Climb-Descent ...............................................................................................

7A-140

Section VIII

Fuel Flow.......................................................................................................

7A-143

Section IX

Airspeed System Characteristics...................................................................

7A-145

Section X

Special Mission Performance........................................................................

7A-148

NORMAL PROCEDURES...........................................................................

8-1

Section I

Mission Planning ...........................................................................................

8-1

Section II

Operating Procedures and Maneuvers ..........................................................

8-3

Section III

Instrument Flight ...........................................................................................

8-17

Section IV

Flight Characteristics.....................................................................................

8-18

CHAPTER 7A

CHAPTER 8

Page

Change 10

PERFORMANCE DATA

TM 1-1520-237-10

TABLE OF CONTENTS (Cont)
Chapter
&
Section
Section V

Page
Adverse Environmental Conditions ..............................................................

8-20

EMERGENCY PROCEDURES ...................................................................

9-1

Section I

Aircraft Systems ............................................................................................

9-1

Section II

Mission Equipment........................................................................................

9-22

APPENDIX A

REFERENCES ..............................................................................................

A-1

APPENDIX B

ABBREVIATIONS AND TERMS...............................................................

B-1

APPENDIX C

KY-100. .........................................................................................................

C-1

INDEX

INDEX ...........................................................................................................

INDEX-1

CHAPTER 9

Change 10

v/(vi Blank)

TM 1-1520-237-10

CHAPTER 1
INTRODUCTION
1.1 GENERAL.
These instructions are for use by the operator. They apply to UH-60A, UH-60L, and EH-60A helicopters.

quired because of multiple emergencies, adverse weather,
terrain, etc. Your flying experience is recognized and therefore, basic flight principles are not included. IT IS REQUIRED THAT THIS MANUAL BE CARRIED IN THE
HELICOPTER AT ALL TIMES.

1.2 WARNINGS, CAUTIONS, AND NOTES.
1.4 APPENDIX A, REFERENCES.
Warnings, cautions, and notes are used to emphasize
important and critical instructions and are used for the following conditions:

WARNING
An operating procedure, practice, etc.,
which, if not correctly followed, could result in personal injury or loss of life.

Appendix A is a listing of official publications cited
within the manual applicable to and available for flight
crews, and fault isolation/trouble references.
1.5 APPENDIX B, ABBREVIATIONS, AND TERMS.
Abbreviations listed are to be used to clarify the text in
this manual only. Do not use them as standard abbreviations.
1.6 INDEX.

CAUTION

An operating procedure, practice, etc.,
which, if not strictly observed, could result in damage to or destruction of equipment.
NOTE
An operating procedure, condition, etc.,
which it is essential to highlight.
1.3 DESCRIPTION.
This manual contains the complete operating instructions
and procedures for UH-60A, UH-60L, and EH-60A helicopters. The primary mission of this helicopter is that of
tactical transport of troops, medical evacuation, cargo, and
reconnaissance within the capabilities of the helicopter. The
observance of limitations, performance, and weight and
balance data provided is mandatory. The observance of
procedures is mandatory except when modification is re-

The index lists, in alphabetical order, every titled paragraph, figure, and table contained in this manual. Chapter 7
performance data has an additional index within the chapter.
1.7 ARMY AVIATION SAFETY PROGRAM.
Reports necessary to comply with the safety program are
prescribed in AR 385-40.
1.8 DESTRUCTION OF ARMY MATERIEL TO PREVENT ENEMY USE.
For information concerning destruction of Army materiel to prevent enemy use, refer to TM 750-244-1-5.
1.9 FORMS AND RECORDS.
Army aviators flight record and aircraft inspection and
maintenance records which are to be used by crewmembers
are prescribed in DA PAM 738-751 and TM 55-1500-34223.

Change 6

1-1

TM 1-1520-237-10

1.10 EXPLANATION OF CHANGE SYMBOLS.
Changes, except as noted below, to the text and tables,
including new material on added pages, are indicated by a
vertical line in the outer margin extending close to the entire area of the material affected: exception; pages with
emergency markings, which consist of black diagonal lines
around three edges, may have the vertical line or change
symbol placed along the inner margin. Symbols show current changes only. A vertical line alongside the title is used
to denote a change to an illustration. However, a vertical
line in the outer margin, is utilized when there have been
extensive changes made to an illustration. Change symbols
are not used to indicate changes in the following:

DESIGNATOR
SYMBOL
EH

EH-60A
information.

Aircraft with NVG lighting.

Aircraft with External Stores
Support Systems.

700

UH-60A, EH-60A aircraft
equipped with T700-GE-700
engines.

701C

UH-60L aircraft equipped
with T700-GE-701C
engines.

a. Introductory material.
b. Indexes and tabular data where the change cannot be
identified.

1.11 SERIES AND EFFECTIVITY CODES.

Aircraft
installed.

UH-60A,
information.

UH-60L

UH−60A

UH-60A
information.

peculiar

UH−60L

UH-60L
information.

peculiar

Change 4

volcano

GPS

Aircraft
with
positioning system
installed.

ERFS

Aircraft
with
Extended
Range Fuel System.

AFMS

Aircraft with Auxiliary Fuel
Management System.

global
(GPS)

will be used throughout this manual to
This symbol
designate information applicable to the high drag configuration described in Chapters 7 and 7A.
1.13 PLACARDED AIRCRAFT SYMBOL.
will be used throughout this manual to
This symbol
designate applicability to helicopters which have torque
placard limitations.
1.14 USE OF WORDS SHALL, SHOULD, AND MAY.

APPLICATION

1-2

with

1.12 HIGH DRAG SYMBOL.

Designator symbols listed below, are used to show limited effectivity of airframe information material in conjunction with text content, paragraph titles, and illustrations.
Designators may be used to indicate proper effectivity, unless the material applies to all models and configuration
within the manual. Designator symbols precede procedural
steps in Chapters 8 and 9. If the material applies to all
series and configurations, no designator symbol will be
used.
DESIGNATOR
SYMBOL

peculiar

VOL

c. Blank space resulting from the deletion of text, an
illustration, or a table.
d. Correction of minor inaccuracies, such as spelling,
punctuation, relocation of material, etc., unless such correction changes the meaning of instructive information and
procedures.

APPLICATION

Within this technical manual the word shall is used to
indicate a mandatory requirement. The word should is used
to indicate a nonmandatory but preferred method of accomplishment. The word may is used to indicate an acceptable
method of accomplishment.

TM 1-1520-237-10

CHAPTER 2
AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION
Section I AIRCRAFT
2.1 GENERAL.

2.3 UH-60L.

This chapter describes the UH-60A, UH-60L, and EH60A helicopter’s systems and flight controls. The functioning of electrical and mechanical components is simplified
where more detailed knowledge is not necessary.

The UH-60L helicopter is the same as the UH-60A helicopter except engines T700-GE-701C replace T700-GE700. The main transmission is replaced by an improved
durability gearbox (IDGB).

2.2 UH-60A.

2.4 EH-60A.

UH−60A

The UH-60A (BLACK HAWK) (Figure 2-1) is a twin
turbine engine, single rotor, semimonocoque fuselage, rotary wing helicopter. Primary mission capability of the helicopter is tactical transport of troops, supplies and equipment. Secondary missions include training, mobilization,
development of new and improved concepts, and support of
disaster relief. The main rotor system has four blades made
of titanium/fiberglass. The drive train consists of a main
transmission, intermediate gear box and tail rotor gear box
with interconnecting shafts. The propulsion system has two
T700-GE-700 engines operating in parallel. The nonretractable landing gear consists of the main landing gear and a
tailwheel. The armament consists of two 7.62 mm machineguns, one on each side of the helicopter in the forward
cabin. Detailed descriptions of these systems are given in
these chapters. For additional weight information, refer to
Chapters 5, 6, and 7. Kit installations for the helicopter
consist of range extension tanks, rescue hoist, medical
evacuation, infrared suppression, blade anti-icing/deicing,
blackout devices, snow skis, winterization and static/
rappelling kit. Refer to this chapter and Chapter 4 for kit
descriptions.

UH−60L

The EH-60A helicopter is a modified UH-60A (Figure
2-1) with a crew of four. The Mission equipment consists
of electronic systems with modifications that will ensure
that the mission requirements are met. The EH-60A system
includes air conditioning, helicopter survivability equipment, and avionics equipment. An electronics compartment
within the transition section is used for avionics equipment.
The compartment can be entered from the right side of the
helicopter. The mission systems employ two operators: The
DF (ESM) operator controlling the electronics surveillance
functions, and the electronics countermeasure (ECM) operator controlling the active countermeasure functions. The
EH-60A can operate independently or in conjunction with
up to two additional, similarly equipped, aircraft. When
operating in the multisystem mode, secured air-to-air communications are provided for automatic tasking between
aircraft. Secured air-to-ground communications are also
provided for voice reporting purposes.

Change 9

2-1

TM 1-1520-237-10

2.5 DIMENSIONS.
Principal dimensions of the helicopter are based on the
cyclic stick and tail rotor pedals being centered and the
collective stick being in its lowest position. All dimensions
are approximate and they are as shown on Figure 2-2.
2.6 TURNING RADIUS AND GROUND CLEARANCE.

WARNING
Main rotor clearance in Figure 2-3 is
shown with cyclic centered and level
ground. Cyclic displacement or sloping
terrain may cause rotor blade clearance
to be significantly less.
For information on turning radius and ground clearance,
see Figure 2-3.
2.7 COMPARTMENT DIAGRAM.
2.7.1 Compartment Diagram. UH The fuselage is divided into two main compartments, the cockpit and cabin.
The cockpit (Figure 2-4) is at the front of the helicopter
with the pilots sitting in parallel, each with a set of flight
controls and instruments. Operation of electrical controls is
shared by both. The cabin compartment contains space for
crew chief seating, troop seating, litter installation and
cargo. Restraint of cargo is by tiedown rings installed in the
floor. Two stowage compartments (Figure 6-11), at the rear
of the cabin over the main fuel tanks, are for flyaway equipment. The equipment storage compartments are reached
from inside the cabin. A gust lock control, APU accumulator handpump and pressure gage, and APU ESU are also
installed (Figure 2-5).
2.7.2 Compartment Diagram. EH A fixed observer
seat is installed to allow observation of either operator position (Figure 2-6). Floor attachments are provided for securing rack mounts and seats. Blackout curtains may be
used to eliminate any light intrusion into the cockpit during
night operations, or any glare on the operator’s console
during day operations. Blackout curtains may be used between cockpit and cabin during NVG operations.
2.8 UPPER AND LOWER CONSOLES.
All cockpit electrical controls are on the upper and lower
consoles and instrument panel. The upper console (Figure
2-7), overhead between pilot and copilot, contains engine
controls, fire emergency controls, heater and windshield
2-2

Change 10

wiper controls, internal and external light controls, electrical systems and miscellaneous helicopter system controls.
The rear portion of the upper panel contains the dc essential
bus circuit breaker panels. The lower console (Figure 2-8)
next to the base of the instrument panel and extending
through the cockpit between the pilot and copilot, is easily
reached by either pilot. The console is arranged with communication panels, navigational panels and flight attitude/
stability controls. The rear part of the console houses the
battery bus and battery utility bus circuit breaker panels,
and parking brake handle.
2.9 LANDING GEAR SYSTEM.
The helicopter has a nonretractable landing gear consisting of two main gear assemblies and a tailwheel assembly.
The landing gear permits helicopter takeoffs and landings
on slopes in any direction. The system incorporates a jack
and kneel feature that permits manual raising or lowering
of the fuselage for air transportability. A landing gear
weight-on-wheels (WOW) switch is installed on the left
landing gear to control operation of selected systems (Table
2-1). The switch is deactivated when the weight of the helicopter is on the landing gear. On helicopters equipped
with ESSS fixed provisions, a WOW switch is also installed on the right landing gear drag beam to provide ac
underfrequency cutout and external stores jettison. The left
WOW switch provides all other WOW functions as without
ESSS provisions and the EMER JETT ALL capabilities.
See Table 2-1 for reference.
2.9.1 Main Landing Gear. The main landing gear is
mounted on each side of the helicopter forward of center of
gravity (Figure 2-1). Each individual landing gear has a
single wheel, a drag beam, and a two-stage oleo shock strut.
The lower stage will absorb energy from landings up to 10
feet-per-second (fps). Above 10 fps the upper stage and
lower stage combine to absorb loads up to 39 fps (about
11.25 Gs).
2.9.2 Wheel Brake System. Main landing gear wheels
have disc hydraulic brakes. The self-contained selfadjusting system is operated by the pilot’s and copilot’s tail
rotor pedals. The brakes have a visual brake puck wear
indicator. Each wheel brake consists of two steel rotating
discs, brake pucks and a housing that contains the hydraulic
pistons. The parking brake handle, marked PARKING
BRAKE, is on the right side of the lower console (Figure
2-8). A hand-operated parking brake handle allows brakes
to be locked by either pilot or copilot after brake pressure is
applied. The parking brakes are applied by pressing the toe
brake pedals, pulling the parking brake handle to its fully
extended position, and then releasing the toe brakes while
holding the handle out. An advisory light will go on, indi-

TM 1-1520-237-10

18 17

1.
2.
3.
4.
5.
6.
7.
EH 8.
9.
10.
EH 11.
12.
13.

12 11 10

PITOT CUTTER
BACK HYDRAULIC PUMP
NO. 1 HYDRAULIC PUMP AND NO.1 GENERATOR
UPPER (ROTOR PYLON) CUTTER
INFRARED COUNTERMEASURE TRANSMITTER
AFT MAINTENANCE LIGHT RECEPTACLE
TAIL LANDING GEAR DEFLECTOR
FLARE DISPENSER
CHAFF DIPENSER
APU EXHAUST PORT
COOLING AIR INLET PORT
PNEUMATIC PORT
PRESSURE AND CLOSED CIRCUIT REFUELING PORTS

14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.

NO. 1 ENGINE
MAIN LANDING GEAR DEFLECTOR / CUTTER
LANDING GEAR JOINT DEFLECTOR
STEP AND EXTENSION DEFLECTOR
DOOR HINGE DEFLECTOR
RIGHT POSITION LIGHT (GREEN)
FIRE EXTINGUISHER BOTTLES
FORMATION LIGHTS
TAIL POSITION LIGHT (WHITE)
APU
LEFT POSITION LIGHT (RED)
PITOT TUBES

ON HELICOPTERS EQUIPPED WITH WIRE STRIKE PROTECTION SYSTEM

AA0403_1A
SA

Figure 2-1. General Arrangement (Sheet 1 of 2)

2-3

TM 1-1520-237-10

40
39

35 34 33

26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
EH 36.
37.

UPPER ANTICOLLISION LIGHT
TAIL DRIVE SHAFT
NO. 2 HYDRAULIC PUMP AND NO. 2 GENERATOR
PYLON CUTTER
HEATER AIR INTAKE PORT
EXTERNAL ELECTRICAL POWER RECEPTACLE
NO. 2. ENGINE
ICE DETECTOR
AMBIENT SENSE PORT
ENGINE FAIRING / WORK PLATFORM (SAME BOTH SIDES)
CONDENSER EXHAUST / STEP
GRAVITY REFUELING PORT (SAME BOTH SIDES)

EH 38.
EH 39.
40.
41.
42.
43.
44.
45.
46.
47.
48.
49.

AFT AVIONICS COMPARTMENT DOORS
IINS BLOWER INLET FILTER
TAIL PYLON FOLD HINGES
TAIL PYLON SERVICE LADDER (SAME BOTH SIDES)
STABILATOR
ENGINE BAY AREA COOLING AIR INTAKE (SAME BOTH SIDES)
WINDSHIELD POST DEFLECTOR
WINDSHIELD WIPER DEFLECTOR
AVIONICS COMPARTMENT
OAT SENSOR
ICE DETECTOR
PYLON COOLING AIR INTAKE

ON HELICOPTERS EQUIPPED WITH WIRE STRIKE PROTECTION SYSTEM

AA0403_2B
SA

Figure 2-1. General Arrangement (Sheet 2 of 2)
2-4

TM 1-1520-237-10

WIDTH WITH ESSS AND EXTERNAL
EXTENDED RANGE TANKS INSTALLED
21 FEET

FUSELAGE WIDTH WITH
HOVER IR SUPPRESSORS
INSTALLED
9 FEET − 8 INCHES

FUSELAGE WIDTH
7 FEET − 9 INCHES
20O

8 FEET−
9 INCHES
5 FEET
1 INCH
3 FEET
9.5 INCHES
TREAD
8 FEET
10.6 INCHES
MAIN LANDING GEAR
9 FEET − 8.6 INCHES
STABILATOR WIDTH
14 FEET − 4 INCHES

TAIL ROTOR
DIAMETER
11 FEET

12 FEET−
4 INCHES
2.8 INCHES
MAIN ROTOR DIAMETER
53 FEET − 8 INCHES

9 FEET −
5 INCHES

WHEEL BASE 29 FEET
7 FEET −
7 INCHES

1 FOOT −
7 INCHES

6 FEET −
6 INCHES
LENGTH − ROTORS AND PYLON FOLDED 41 FEET − 4 INCHES

FUSELAGE LENGTH 50 FEET − 7.5 INCHES
OVERALL LENGTH 64 FEET − 10 INCHES
AA0514B
SA

Figure 2-2. Principal Dimensions
2-5

TM 1-1520-237-10

TURNING
RADIUS
41 FEET
7.7 INCHES

O
* TAIL ROTOR IS CANTED 20 . UPPER
TIP PATH PLANE IS 16 FEET 10 INCHES
ABOVE GROUND LEVEL

12 FEET
4 INCHES
9 FEET
5 INCHES
ROTOR
TURNING

16 FEET *
10 INCHES
7 FEET
7 INCHES
ROTOR
STATIONARY
12 FEET
1 INCH

6 FEET
6 INCHES
11 FEET
4 INCHES

WHEELBASE 29 FEET

Figure 2-3. Turning Radius and Clearance
2-6

12 FEET
5 INCHES

AA0402
SA

TM 1-1520-237-10

25
24
1
23
2
3

3
4

22
21
6
20
5
7

12
9

13
ST
LI
K AP
EC M
CH A & GE
T A
DATOW
S

CH
DA ECK
T
ST A & LIST
OW MA
AG P
E

1. UPPER CONSOLE
2. PILOT’S COCKPIT UTILITY LIGHT
3. FREE−AIR TEMPERATURE GAGE (ON
HELICOPTERS WITH HEATED CENTER
WINDSHIELD)
4. NO. 2 ENGINE FUEL SELECTOR LEVER
5. NO. 2 ENGINE OFF / FIRE T−HANDLE
6. NO. 2 ENGINE POWER CONTROL LEVER
7. WINDSHIELD WIPER
8. INSTRUMENT PANEL GLARE SHIELD

17
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.

INSTRUMENT PANEL
VENT / DEFOGGER
ASHTRAY
PEDAL ADJUST LEVER
MAP / DATA CASE
CABIN DOME LIGHTS DIMMER
CHAFF RELEASE SWITCH
PARKING BRAKE LEVER
FUEL BOOST PUMP PANEL
LOWER CONSOLE UTILITY LIGHT

EH
EH

19.
20.
21.
22.
23.

STANDBY (MAGNETIC COMPASS)
NO. 1 ENGINE POWER CONTROL LEVER
NO. 1 ENGINE OFF / FIRE T−HANDLE
NO. 1 ENGINE FUEL SELECTOR LEVER
AC ESNTL BUS CIRCUIT BREAKER
PANEL
24. COPILOT’S COCKPIT UTILITY LIGHT
25. FREE−AIR TEMPERATURE GAGE (ON
HELICOPTERS WITHOUT HEATED
AB0821
CENTER WINDSHIELD)
SA

Figure 2-4. Cockpit Diagram (Sheet 1 of 2)

Change 4

2-7

TM 1-1520-237-10

28
28

30
30

31
32

ST
LI
K AP
EC M
CH A & GE
T A
DATOW
S

CH
DA ECK
T
ST A & LIST
OW MA
AG P
E

38
34
45

36
44
26.
27.
28.
29.
30.
31.
32.
33.

COCKPIT FLOODLIGHT CONTROL
UPPER CONSOLE
MASTER WARNING PANEL
SLIDING WINDOW
COCKPIT DOOR EMERGENCY RELEASE
CYCLIC STICK
DIRECTIONAL CONTROL PEDALS
PILOT’S SEAT

34. CREW CHIEF / GUNNER ICS CONTROL
PANEL
35. CREW CHIEF AMMUNITION / GRENADE
STOWAGE COMPARTMENT
36. STOWAGE BAG
37. COLLECTIVE STICK FRICTION CONTROL
38. COLLECTIVE STICK GRIP
39. ENGINE IGNITION KEYLOCK
40. LOWER CONSOLE

38 37
41. BATTERY / BATTERY UTILITY BUS
CIRCUIT BREAKER PANEL
42. FIRE EXTINGUISHER
43. GUNNER’S ICS CONTROL PANEL
44. FIRST AID KIT
45. GUNNER’S AMMUNITION / GRENADE
46. COPILOT’S SIDE LOWER CONSOLE
AUXILIARY FUEL MANAGEMENT PANEL
47.
AFMS
AB0822
SA

Figure 2-4. Cockpit Diagram (Sheet 2 of 2)

2-8

Change 4

TM 1-1520-237-10

E
CABIN DOME
LIGHTS (THREE)

D
TROOP COMMANDER’S
ANTENNA COAX

C
B
ON HELICOPTERS EQUIPPED
WITH AUXILIARY CABIN HEATER

GUST LOCK
RELEASE
BUTTON

STA
378.50

STA
349.50

GUST LOCK
HANDLE

GUST LOCK CONTROL
STA
332.50

ACCUMULATOR
PRESSURE GAGE

HEATER AIR
INLET PORT

ACCUMULATOR
HAND PUMP

ACCUMULATOR
MANUAL
START
VALVE

HEATER TEMPERATURE
CONTROL

50 85

ACCUMULATOR
PISTON POSITION
INDICATOR

ON HELICOPTERS EQUIPPED WITH
AUXILIARY CABIN HEATER

APU ACCUMULATOR
(LOOKING UP)

(LOCATED BELOW LEFT GUNNER’S WINDOW)

AA0323_1B
SA

Figure 2-5. Cabin Interior (Sheet 1 of 2)
cating PARKING BRAKE ON. Pressing either pilot or
copilot left brake pedal will release the parking brakes, the
handle will return to the off position and the advisory light
will go off. Power for the advisory light comes from the
No. 1 dc primary bus through a circuit breaker marked
LIGHTS ADVSY.
2.9.3 Tail Landing Gear. The tail landing gear (Figure
2-1) is below the rear section of the tail cone. It has a
two-stage oleo shock strut, tailwheel lock system fork assembly, yoke assembly, and a wheel and tire. The fork
assembly is the attachment point for the tailwheel and allows the wheel to swivel 360°. The tailwheel can be locked
in a trail position by a TAILWHEEL switch in the cockpit

indicating LOCK or UNLK (Figure 2-8). The fork is
locked by an electrical actuator through a bellcrank and
locking pin. When the pin is extended, the switch will indicate LOCK. When the pin is retracted, the switch will
indicate UNLK. Power to operate the locking system is by
the dc essential bus through a circuit breaker marked
TAILWHEEL LOCK.
2.10 INSTRUMENT PANEL.
2.10.1 Instrument Panel. UH Engine and dual flight
instruments are on the one-piece instrument panel (Figure
2-9). The panel is tilted back 30°. The master warning panels are mounted on the upper instrument panel below the

2-9

TM 1-1520-237-10

FAULT INDICATION

FAULTS

T−62T−40−1
1

BITE #
2 3 4

DECODED BITE INFORMATION
ON

RESET
START FUEL VALVE & EXCITER SIGNAL OUT (5%)
MAIN FUEL VALVE SIGNAL OUT (14%)
START FUEL VALVE & EXCITER SIGNAL OFF (70%)
90% RPM SWITCH ON
READY FOR SERVICE (90% + 1.5 SEC)
PROCESSOR BOARD FAILURE
SENSOR/DATA BOARD FAILURE
OVERTEMPERATURE
OVERSPEED
UNDERSPEED
FAIL TO START
LOW OIL PRESSURE
HIGH OIL TEMPERATURE (WARNING)
FAIL TO LIGHT (NO DATA)
SHORTED THERMOCOUPLE PROBE (WARNING)
OPEN THERMOCOUPLE
PROCESSOR SEQUENCE FAIL
NO DATA

START
SEQUENCE

OPERATION

OFF
C
O
M
M

VOL

AUX

ICS

NAV

OFF
C
O
N
T
HOT MIKE

OFF

TROOP COMMANDER’S ICS CONTROL

APU ELECTRONIC SEQUENCE UNIT FAULT INDICATION
(ON HELICOPTERS EQUIPPED WITH T−62T−40−1 APU)

FAULT INDICATION

FAULTS

GTC−P36−150
1
START
SEQUENCE

OPERATION

BITE #
2 3 4

DECODED BITE INFORMATION
RESET (START INITIATED)
FUEL VALVE AND IGNITION SIGNAL ON (5%)
START VALVE SIGNAL OFF (70%)
IGNITION SIGNAL OFF (95%)
READY FOR SERVICE (LOSS OF DC POWER)
ESU FAILURE
A/C START SYSTEM FAILURE
OVERTEMPERATURE
OVERSPEED
UNDERSPEED
FAIL TO START
LOW OIL PRESSURE
OIL PRESSURE SWITCH FAILED
THERMOCOUPLE FAILED
MONOPOLE FAILED
FUEL SOLENOID FAILED
FUEL TORQUE MOTOR FAILED
IGNITION UNIT FAILED
HOT SENSOR FAILED
NO DATA

PUSH−TO−TALK
SWITCH

TROOP COMMANDER’S HANDSET

APU ELECTRONIC SEQUENCE UNIT FAULT INDICATION

AA0323_2A

(ON HELICOPTERS EQUIPPED WITH GTC−P36−150 APU)

Figure 2-5. Cabin Interior (Sheet 2 of 2)
glare shield, to inform the pilot of conditions that require
immediate action.
2.10.2 Instrument Panel. EH The instrument panel of
the EH-60A is as shown on Figure 2-9. Refer to Chapter 3
for description and operation of systems switch panels and
Chapter 4 for BDHI, CREW CALL switch, FLARE
switch and ECM ANTENNA switch and countermeasure
set ALQ-156.
2.10.3 Vertical Instrument Display System (VIDS).
The VIDS (Figure 2-9) consists of a vertical strip central
display unit (CDU), two vertical strip pilot display units
(PDU), and two signal data converters (SDC). Those read-

2-10

ings are shown by ascending and descending columns of
multicolored lights (red, yellow, and green) measured
against vertical scales which operate in this manner: the
segments will light in normal progression and remain on as
the received signal level increases. Those scales will go off
in normal progression as the received signal level decreases.
Scales with red-coded and/or amber-coded segments below
green-coded segments operate in this manner: When the
received signal level is zero or bottom scale, the segments
will light in normal progression and will remain on. When
the first segment above the red or amber range goes on, all
red-coded or amber-coded segments will go off. These segments will remain off until the received signal level indicates a reading at or within the red or amber range. At that

TM 1-1520-237-10

PILOT

COPILOT

ECM CONSOLE

DF CONSOLE

ECM OPERATOR SEAT

DF OPERATOR SEAT

ECM EQUIPMENT RACK MISSION INTERFACE PANEL

OBSERVER SEAT

Figure 2-6. Cabin Mission Equipment Arrangement
time all red-coded or amber-coded segments will go on and
the scale display will either go on or go off in normal progression, depending upon the received signal level. The
CDU and PDUs contain photocells that automatically adjust lighting of the indicators with respect to ambient light.
If any one of the three photocells should fail, the lights on
the vertical scales of the PDUs or CDU may not be at the
optimum brightness for the ambient conditions. The DIM
knob on the CDU contains an override capability which
allows the pilot to manually set the display light level. The
SDCs receive parameter data from the No. 1 and No. 2
engines, transmission, and fuel system; provides processing
and transmits the resulting signal data to the instrument

DF EQUIPMENT RACK

AA0401
SA

display. The No. 1 engine instruments on the CDU and
copilot’s PDU, receive signal data from the No. 1 SDC
(CHAN 1). The No. 2 engine and main transmission instruments on the CDU and pilot’s PDU, receives signal data
from the No. 2 SDC (CHAN 2). If either SDC fails, the
corresponding CHAN 1 or 2 light will go on, and it is
likely the pilot’s or copilot’s PDU and the corresponding
instruments will fail. Failure of a lamp power supply within
an SDC will cause every second display light on the CDU
to go off. Both SDCs receive % RPM 1 and 2, % RPM R
and % TRQ information from both engines. Therefore if
one SDC fails only one PDU will provide % RPM 1 and 2
and % TRQ for both engines.

Change 9

2-11

TM 1-1520-237-10

Table 2-1. Weight-On-Wheels Functions
WOW SWITCH FUNCTION

ON GROUND

IN FLIGHT

Backup Pump Automatic Operation

Disabled (Except when APU
accumulator is low)

Enabled

Hydraulic Leak Test System

Enabled

Disabled

Backup Pump Thermal Switch

Enabled

Disabled

Low % RPM R Audio Warning

Disabled

Enabled

SAS/FPS Computer

Degraded

Enabled

Generator Underfrequency Protection

Enabled

Disabled

IFF Mode 4 Operation

Disabled Automatic Zeroize

Enabled Automatic Zeroize

Disabled

Enabled

Deleted

Enabled

Disabled

5 minute delay before continuous BIT
(CBIT) monitors attitude sensor after AFMP
power up AFMS

Enabled

Disabled

AFMP Power up BIT (PBIT) and Initiated
BIT (IBIT) AFMS

Enabled

Disabled

External Stores Jettison

Deleted
AUX FUEL INCR/DECR Switch

2.10.4 Central Display Unit (CDU). The CDU (Figure
2-9) contains instruments that display fuel quantity, transmission oil temperature and pressure, engine oil temperature and pressure, turbine gas temperature (TGT), and gas
generator speed (Ng) readings. Those readings are shown
by ascending and descending columns of multicolored
lights (red, yellow, and green) measured against vertical
scales. If the instrument contains low range turnoff (red or
yellow lights below green lights) they will go off when the
system is operating within the normal range (green). If the
instrument contains yellow or red lights above the green
range, the green as well as the yellow or red will stay on
when operating above the green range. The operating
ranges for the different instruments are shown in Figures
5-1, 5-2, and 5-3. Digital readouts are also installed on the
TOTAL FUEL quantity, TGT, and Ng gages.
2.10.4.1 Lamp Test System. The lamp test provides a
means of electrically checking all CDU scale lamps, digital
readouts, and % RPM RTR OVERSPEED lights on the
PDUs. When the PUSH TO TEST switch on the CDU is
pressed, all CDU scale lamps should light, digital readouts
should display 888, and three RTR OVERSPEED lights
on the PDUs should be on.

2-12

Change 10

2.10.4.2 Dim Control. The DIM control allows the pilot
to set a desired display light level of the CDU and PDUs in
accordance with the ambient light, or override the auto-dim
sensors. If the auto-dim circuitry should fail or malfunction,
turn the DIM control fully clockwise to regain illumination
of the CDU and PDUs.
2.10.4.3 CDU and PDU Digital Control. An ON, OFF
DIGITS control switch is on the CDU (Figure 2-9) to turn
on or off the digital readout displays on the CDU and PDUs.
If a digital processor fails, all digital displays will go off.
2.10.5 Pilot’s Display Unit (PDU). The PDU (Figure
2-9) displays to the pilot engine power turbine speed (%
RPM 1 and 2), rotor speed (% RPM R), and torque (%
TRQ). Readings are shown by ascending and descending
columns of multicolored lights (red, yellow, and green)
measured against vertical scales. A TEST switch provides
a means of electrically checking all PDU scale lamps and
digital readouts. When the TEST switch is pressed, all PDU
scale lamps should light and digital readouts should display
188. The % RPM indicators contain low range turnoff below the normal operating range. Three overspeed lights at
the top will go on from left to right when a corresponding

TM 1-1520-237-10

rotor speed of 127%, 137%, and 142% is reached. Once a
light is turned on, a latch prevents it from going off until
reset by maintenance. Power for the PDUs is from No. 1
and No. 2 ac and dc primary buses through circuit breakers

marked NO. 1 AC INST/NO. 1 DC INST and NO. 2 AC
INST/NO. 2 DC INST respectively. See Figures 5-1, 5-2,
and 5-3 for instrument markings.

Change 8

2-12.1/(2-12.2 Blank)

TM 1-1520-237-10

WHITE

28V #387

O
F
F

OPEN

BLUE

SPARE
LAMPS

DC ESNTL BUS

E
ICS

NO. 1 VOR / ILS CHIP
2

PILOT COPILOT VHF FM

COMM SCTY SET UHF
NO. 1 FM UHF AM AM
2

ESSS
JTSN

7.5

5
DET

CARGO HOOK

FUEL
DUMP

EMERG REL
TEST
NORM
O
P
E
N
SHORT

CONTR OUTBD

PNL

ARMING
SAFE

ARMED

ALL

CONTR

INBD

3
2

PWR

EMER

RATE

ENG

SENSE

SPLY

BRT

NO. 1
TAIL
ENG WHEEL

SAS

SEC

CONTR SRCH

LOCK

PNL

PWR

CONTR

LIGHTS

APU

EXT PWR
RESET

BATT
O
F
F

NO. 1 ENG OVSP
TEST A
TEST B

R
O E
F S
F E
T

APU
TEST

R
O E
F S
F E
T

FIRE DETR TEST
OPER

NO. 1
TEST

R
O E
F S
F E
T

NO. 2
TEST

WINDSHIELD
WIPER
HEATER
MED

VENT
BLOWER

OFF
LOW

PARK

O
F
F

NO. 2 ENG OVSP
TEST A
TEST B

OFF
CONSOLE LT

CPLT FLT
INST LTS

LIGHTED
SWITCHES

BOOST START

1
OFF

OFF

O
F
F

BATT
BUS

GENERATORS
GLARESHIELD
LIGHTS

ESNTL
DC

FIRE DET
NO.1
NO.2

7.5

FIRE EXTGH

OFF

CARGO PILOT
HOOK TURN

7.5

CONTR SHEAR

FORMATION LT
5
4

STAB

APU

CAUT / BACKUP HOIST ESSS
HYD CABLE JTSN
ADVSY

7.5

DC ESNTL BUS

CONTR
CKPT

LOWER

UPPER

OFF

BRT

NAV LTS
N
O
R
M
IR

BRT

OFF
CARGO
HOOK LT

CABIN
DOME LT
WHITE

FUEL PUMP
APU BOOST
O
F
F
FUEL PRIME

FIRE EXTGH
RESERVE
O
F
F

AIR SOURCE
HEAT / START
ENGINE
O
F
F
APU

OFF

O
F
F

BRT

FLASH

BLUE

ANTICOLLISION
LIGHTS
DAY
UPPER
O
F
F

LOWER

NIGHT

BRT

ENG ANTI−ICE
NO. 1
NO. 2
O
F
F

O
F
F

ON
BACKUP
HYD PUMP
OFF
A
U
T
O
ON

PILOT FLT

NON FLT

MAIN

O
F
F

B
O
T
H

BRT

OFF
INSTR LT

OFF
POSITION
LIGHTS
STEADY
DIM

BRT

HYD
LEAK TEST
RESET
N
O
R
M
TEST

WINDSHIELD
COPILOT
O
F
F

BRT

OFF

ANTI−ICE
PILOT
O
F
F

ON
PITOT
HEAT

O
F
F

AA0364_1A
SA

Figure 2-7. Upper Console (Sheet 1 of 2)
2-13

TM 1-1520-237-10

DC ESNTL BUS

ECS
AIR COND
COOL

TEMP CONT

O
F
F

Q/F
PWR

HTR
O
F
F

O
F
F

FAN

PILOT
TURN

STAB
7.5

RATE

ENG

SENSE

SPLY

NO. 1
TAIL
ENG WHEEL

SAS

APU
CONTR

BATT
BUS

PWR

WARM

COOL

ESNTL
DC

FIRE DET
NO.1
NO.2

LIGHTS
SEC

CONTR SRCH

FIRE EXTGH
5

BOOST START LOCK

PNL

PWR

CONTR

APU

C
DC ESNTL BUS
ICS
2

ESSS
JTSN

NO. 1 VOR / ILS CHIP
5

PILOT COPILOT VHF FM

7.5

DET

OUTBD

EH
WINDSHIELD
WIPER

E
COMM SCTY SET UHF
NO. 1 FM UHF AM AM
2

LOW

ESSS
JTSN

CAUT / BACKUP
HYD
ADVSY

7.5

PNL

CONTR

INBD

HEATER
MED

VENT
BLOWER

OFF
PARK

O
F
F

OFF

CONSOLE LT
UPPER

OFF

LOWER

BRT

OFF
INST LT

NON FLT

OFF

O
F
F

BRT

OFF

O
F
F

WINDSHIELD ANTI−ICE
COPILOT
CTR
PILOT
O
F
F

PITOT
HEAT

O
F
F

WINDSHIELD ANTI−ICE
COPILOT CTR PILOT
O
F
F

BRT

ENG ANTI−ICE
NO. 1
NO. 2

O
F
F

PILOT FLT

O
F
F

(ON HELICOPTERS EQUIPPED WITH
HEATED CENTER WINDSHIELD)

AA0364_2A
SA

Figure 2-7. Upper Console (Sheet 2 of 2)
2-14

TM 1-1520-237-10

STORES JETTISON

EMER
JETT
ALL

INBD

OUTBD

BOTH

OFF

COMPASS

M
I
S
C

JETT

TAIL SERVO
NORMAL

S
W

FUEL
IND
TEST

GYRO
ERECT

TAIL
WHEEL

R
C
V
R

DIM

COMP
OFF

ANT

G
P
S
/
D
P
L
R

LOOP
LOOP

PP
GS/TK
NAV M

AUX

DISPLAY

RIGHT

GHI
3

TGT
STR

JKL
4

MNO
5

PQR
6

INC
(+)

STU
7

VWX
8

YZ*
9

CLR

#
0

ENT

1
2

3
4

5
6

6
FILL

POWER

CHAN
1

LAT /
LONG
GPS
LDG

OFF

DEC
(−)

MODE

OFF

LTR

MID

DEF
2

I
Z
E

MGRS

LAMP
TEST

NAV

LTR

3
2

ROUTE

TEST

N
A
V

LTR

LEFT

ABC
1

WP
TGT

RV
DELAY
Z
E
R
O

DATUM

KYBD

DIST / BRG
TIME

XTK/TKC
KEY
WIND−UTC
DATA

MODE
OP

KY
58
R
C
U

VOICE

PLAIN
C / RAD

MAL

TEST

AUDIO

PUSH TO
SET

TGT
STR

1 7 : BANDO 0 3 0MG 9 1
GP S : M NA V : C
GS : 1 1 7 KM / HR
TK : 0 2 5 "

BRT

SYS
STAT

EPE

SLAVED

BACKUP
FREE

KILOCYCLES

T
U
N
E

ALL
FLY TO

A
D
F

(PAGE)

OFF
2

C
O
M
M

1
ICS

VOL

C
O
N
HOT MIKE T

4
5

STABILATOR CONTROL
MAN SLEW
UP

OFF

NAV VOL

AUTO
CONTROL

TEST

O
F
F

MAIN
OFF

R
E
S
E
T

ON
DN

MB VOL

VOL
ADF

OFF
ON
SQUELCH

TONE

AUTO FLIGHT CONTROL

108.00

SAS 1

OFF

VOR / MB
TEST

MB SENS

SAS 2

FPS

BOOST

AUX

R
E
S
E
T

MODE

ADVISORY

CPTR SAS 2

ACCL

CLTV

TRIM

A/S

GYRO

RGYR

R
E
S
E
T

OFF
2

C
O
M
M

C
O
N
HOT MIKE T

ICS

VOL

OFF

POWER ON RESET

OP
LD

PLAIN
C / RAD

MODE
OP

DELAY

POWER

TEST/MON

DI
S

TO T

M−1

M−2

M−C

M−3 / A

IFM RF PWR
NORM HI
LO
OFF

6
1

L E

O
N

FREQ

CLR

H−Ld

ERF
OFST

TIME

HOM

Sto
ENT

SC
FH
FH−M

I
F
F

AUDIO
L
I
G
H
T

O
N

MODE 1

VOL

TEST

DI
S

2
1
2

IDENT

MIC

O
U
T

5
6

6
FILL

POWER

IFM RF PWR
NORM HI
LO
6

MAN

OFF

CUE

CLR

H−Ld

L E

FREQ

ERF
OFST

TIME

MODE
FUNCTION
RXMT
LD
SQ OFF
SQ ON
LD−V
Z−A
TEST

MODE 3 / A

3
4

PRESET
2 3 4
5
1

OFF

C
O
M
M

T
O
N
E

ANT

TO T

OUT

0
V
O
L
S
Q
D
I
S

KIT

REPLY
P RES

5
8

RV
DELAY
Z
E
R
O
I
Z
E

STATUS

MODE 4
H
D
OL

4
7

KY
58
R
C
U

OFF

ALT

MODE
FUNCTION
RXMT
LD
SQ OFF
SQ ON
LD−V
Z−A
TEST
OFF
STOW

OUT
CODE

ZERO

CUE

D
I
V

BOT
RAD
TEST

PRESET
2 3 4
5

MASTER

A
N
T

G
O

TEST

O
N

1
MAN

TOP

N
O

TO T

FILL

NORM ST

G
O

TEST

2
1

5
6

P RES

3
4

EM
E

1
2

P RES

Z
E
R
O
I
Z
E

NAV

FAILURE

KY
58
R
C
U

OFF

PLAIN
C / RAD

TRIM

PRESET
MANUAL
GUARD

BOTH

U
H
F

STOW

HOM

Sto
ENT

SC
FH
FH−M

VOL

CURSOR

EMER

PRESET

MAN LOAD
PRE

PWR

SELF
ON
OFF

9
TR

DSCRM

+
TEST

ON
OFF

AUDIO

OFF

FLARE

DIM

E
ES
S TO T

VALUE

SILENT

T/R

STBY

ARM

MAN

R
I F
P I
P R
L E
E

DATA

VOL

NO. 1
PUMP

OFF

CONTR
INST

+
MODE

ON
ADJ

1−4
INC
SEL
DCLT

B
U CKPT
S

PGM

5
GEN
CONTR

ALT/P/R

5
FIRE
DET

CHAFF
DISPENSE

DEC
NXT

5
CONTR
INST

APU

BIT

FIRE

EXTGH

ACK

1−4

T
T

DCLT

CONV EXT PWR BOOST CONTR
U
WARN CONTR
T UTIL
APU
I LTS

P−PGM

BATT &
ESNTL DC FUEL
BATT
WARN PRIME B BUS

SQL

NO. 2
PUMP

PARKING BRAKE
MODE
PLT

CPLT

DSPL POS
BRT D / U

DSPL POS

DIM L / R

L/R

DIM

D / U BRT

FAIL

OFF
FM 1 / UHF

VHF / UHF

FM 1 / VHF

FM 2 / VHF

FM 1 / FM 2

FM 2 / UHF

RADIO RETRANSMISSION

U
AUDIO

BAT
2

MODE
PT
CT

MAN
PWR
OFF

OFL EB Z ALL
(PULL)
RK
CIK

INIT

KY−
100

PRESET
REM
6

1
2 3

PNL
OFF
DSPL
OFF

BRT

FILL

B
U
S

3
2
1

ECCM
EMER

KEY

OFF

ESNTL BUS
AC &
DC

ALE

PGRM

OFF

SAFE

B
50
A
T SPLY
T

PRE
MAN

ZERO
(PULL)

OFF

FUEL BOOST PUMP CONTROL

CP−PGM

DISP
CONT

CHAFF

ARM

AA0385_1E
SA

Figure 2-8. Lower Console (Sheet 1 of 3)

Change 10

2-15

TM 1-1520-237-10

A
KILOCYCLES
CW

80
2

T
U
N
E

A
D
F
R
C
V
R

VOICE

TEST

COMP
OFF

AUDIO

ANT

LOOP
LOOP
R

B
NAV VOL

MB VOL

108.00
OFF

OFF

VOR / MB
TEST

MB SENS
HI
LO

C
FLY TO

BRT

DIM

G
P
S
/
D
P
L
R

EPE

TGT
STR

1 7 : BANDO 0 3 0MG 9 1
GP S : M NA V : C
GS : 1 1 7 KM / HR
TK : 0 2 5 "

PP
GS/TK
NAV M

DIST / BRG
TIME
WP
TGT

XTK/TKC
KEY
WIND−UTC
DATA

MAL

LTR

LEFT

MID

RIGHT

ABC
1

DEF
2

GHI
3

TGT
STR

JKL
4

MNO
5

PQR
6

INC
(+)

STU
7

VWX
8

YZ*
9

DEC
(−)

CLR

#
0

(PAGE)

KYBD

DATUM
ROUTE

DISPLAY
TEST

N
A
V

SYS
STAT

MGRS
LAT /
LONG

LAMP
TEST

GPS
LDG

OFF

MODE

ENT

AA0385_2C

GPS

Figure 2-8. Lower Console (Sheet 2 of 3)

2-16

Change 8

TM 1-1520-237-10

2.11 DOORS AND WINDOWS.
2.11.1 Cockpit Doors. The crew compartment is
reached through two doors, one on each side of the cockpit.
The doors swing outward and are hinged on the forward
side (Figure 2-1). Each door has a window for ventilation.
Installed on the back of each door is a latch handle to allow
unlatching the door from either inside or outside the cockpit. Emergency release handles are on the inside frame of
each door (Figure 9-1). They allow the cockpit doors to be
jettisoned in case of an emergency. There is an emergency
release pull tab on the inside forward portion of each cockpit door window for pilot egress.

2-1). Single-action door latches allow the doors to be
latched in the fully open or fully closed positions. Each of
the two doors incorporate two jettisonable windows, for
emergency exit (Figure 9-1).
2.11.3 Crew Chief/Gunner Windows. The Crew
Chief/Gunner Stations have forward sliding hatch windows,
split vertically into two panels (Figure 2-1). A springloaded security latch is installed on each gunner’s aft window, to prevent the window opening from the outside. The
dead bolt lock requires activation of the security latch lever
from inside the helicopter. Another window latch bar is
actuated to allow the forward window to be moved to a

2.11.2 Troop/Cargo (Cabin) Doors. Aft sliding doors
are on each side of the troop/cargo compartment (Figure

Change 1

2-16.1/(2-16.2 Blank)

TM 1-1520-237-10

COMPASS

M
I
S
C

TAIL SERVO
NORMAL

S
W

COVER

FUEL
IND
TEST

GYRO
ERECT

TAIL
WHEEL

BACKUP
PUSH TO
SET

FREE

PLAIN
C / RAD

KILOCYCLES

T
U
N
E

A
D
F

MODE
OP

LD
KY
58
R
C
U

VOICE

TEST

RV
DELAY

R
I

COMP
OFF

AUDIO

ANT

1
2

LOOP
1

DEST

GHJ

STR

AUX

I
I
N
S

NAV

OFF
2

C
O
M
M

1
ICS

VOL

M
2

KLM

NPQ

UPDT

W
4

5
UVW
S

7
NAV

NORM
DEST

CLR

BRT

USE

OFF
TEST

MAIN
OFF

MB VOL

MAN SLEW
UP

OFF

OFF
MB SENS

SAS 1

SAS 2

PLAIN
C / RAD

BOOST

FAILURE

ADVISORY

DELAY

CPTR SAS 2

ACCL

CLTV

1
2

3
4

5
6

AUX

NAV

OFF
3

C
O
N
HOT MIKE T

ICS

VOL

TRIM

A/S

GYRO

RGYR

OFF
R
E
S
E
T

POWER ON RESET

C
O
M
M

FPS

R
E
S
E
T

RV
Z

OFF
ON
SQUELCH

MODE
OP

TRIM

PRESET
MANUAL
GUARD

OFF

AUTO FLIGHT CONTROL

VOL

TONE

R
E
S
E
T

ADF

AUTO
CONTROL

TEST

O
F
F

VOR / MB
TEST

BOTH

U
H
F

STABILATOR CONTROL
NAV VOL

FACK

CAL

OFF

KY
58
R
C
U

POWER

CHAN

ATTD

108

6
FILL

XYZ

LTR

FAST

C
O
N
HOT MIKE T

DEF

RST
BIT

OFF

ABC

TCN

POS

5
6

STR

INS

3
4

LOOP
MRK

R
C
V
R

SLAVED

POWER

MODE
OP
LD

KY
58
R
C
U

6
FILL

PLAIN
C / RAD

RV
DELAY

O
I

G
O

TO T

TEST

M−3/A

D
I
V

BOT
RAD
TEST

M−C

O
N

MASTER

1
2

A
N
T

NORM ST

M−2

TOP

N
O

OFF
BY

O
N

ALT

CODE
A

PRESET

MAN LOAD
PRE

I
F
F

TEST

OUT

TR
MODE 1

AUDIO
L
I
G
H
T

O
N

OFF

ANT

IFM CONTROL

IDENT

M
DI
S

TO T

MIC

OUT

O
U
T

FLARE

MODE 3 / A

DISP
CONT

ARM

ARM
R
I F
P I
P R
L E
E

NO. 2
PUMP
RADIO RETRANSMISSION
OFF

OFF

ON
NO. 1
PUMP

CHAFF

FUEL BOOST PUMP CONTROL
ON

POWER

MODE 4
H

P RES

EMER

KIT

6
FILL

DIM

C
O
M
M

ZERO

14
T
O
N
E

D
OL

S
Q
D
I
S

5
6

VHF−FM NO. 1
AN/ARC−201

STATUS

OUT
V
O
L

3
4

E
ES
S TO T

M−1

TO T

M
ES

T
ES

P RES

VHF−FM NO. 2
AN/ARC−201

TEST/MON

EM
E

TEST
G
O

FM 1 / FM 2
FM 1 / VHF
FM 1 / UHF

MAN

PGRM

SAFE

FM 2 / UHF
PWR

FM 2 / VHF

SELF

DSCRM

ON
OFF

VHF / UHF

TEST

ON
OFF

AUDIO

OFF
ALQ
162

COVER
AN/ALQ−162

B 50
A
T SPLY
T
B
U
S

5
CONV
WARN
APU

CONTR
INST

FIRE
DET

BATT &
ESNTL DC FUEL BATT
WARN PRIME B BUS

FIRE

A
5
T
T
EXT PWR BOOST CONTR
U
CONTR
T UTIL
I LTS
L

GEN
CONTR

B
U
CKPT
S

5
EXTGH
APU

NO
GO

CW
JAM

ALQ−156

CM
JAM

CM
INOP

PARKING BRAKE

ESNTL BUS
AC &
DC

CW
THRT

ALQ−144

IRCM
INOP

CHAFF
DISPENSE

5
CONTR
INST

AA1304_2B
SA

Figure 2-8. Lower Console (Sheet 3 of 3)

2-17

TM 1-1520-237-10

#1 ENG
OUT

#2 ENG
OUT
MASTER CAUTION
PRESS TO RESET

FIRE

MA WRN

RAD ALT

RADIO CALL
00 0 00

LOW ROTOR
RPM

CAUT/ADVSY

FUEL

XMSN

QTY
LB X 100

NVG DIMMING

DIMMING

#1 FUEL LOW

#1 GEN

#2 GEN

#2 FUEL LOW

#1 ENGINE
PRESS

#1 GEN BRG

#2 GEN BRG

#2 ENGINE
PRESS

TEMP
C X 10

% RPM

250

CMD

CLI MB

200

KNOTS

ABS

% TRQ

TEST RTR OVERSPEED

FT X 100

ATT

130

ALT

150

SET

G
S

FEET

SET
PUSH
TO TEST

OFF

30
40
DN

DI VE

NAV
KIAS
LIMIT
150
100
80
60
45

3 4

BACK
CRS

FM
HOME

BACK
CRS

FM
HOME

NORM
ALTR

PLT
CPLT

NORM
ALTR

ADF
VOR

TURN
RATE

CRS
HDG

VERT
GYRO

BRG
2

2 9 9 0

2
3

#2 ENGINE
OIL TEMP

CHIP
#1 ENGINE

BATT LOW
CHARGE

BATTERY
FAULT

CHIP
#2 ENGINE

#1 FUEL
FLTR BYPASS

GUST
LOCK

PITCH BIAS
FAIL

#2 FUEL
FLTR BYPASS

120

#1 ENGINE
STARTER

#1 OIL
FLTR BYPASS

#2 OIL
FLTR BYPASS

#2 ENGINE
STARTER

110

100

105

100

110

5 4

#1 PRI
SERVO PRESS

#1 HYD
PUMP

#2 HYD
PUMP

#2 PRI
SERVO PRESS

TAIL ROTOR
QUADRANT

IRCM
INOP

* AUX FUEL

#1 TAIL
RTR SERVO

MAIN XMSN
OIL TEMP

INT XMSN
OIL TEMP

TAIL XMSN
OIL TEMP

APU OIL
TEMP HI

BOOST SERVO
OFF

STABILATOR

SAS OFF

8
4
0
−4

2
ON

DIM

OFF
DIGITS

TO TEST
FLT PATH
STAB

IFF

RT PITOT
HEAT

CHIP
INT XMSN

CHIP
TAIL XMSN

CHIP INPUT
MDL−RH

CHIP ACCESS
MDL−LH

CHIP MAIN
MDL SUMP

APU
FAIL

CHIP ACCESS
MDL−RH

MR DE−ICE
FAIL

MR DE−ICE
FAULT

TR DE−ICE
FAIL

ICE
DETECTED

MAIN XMSN
OIL PRESS

#1 RSVR
LOW

#2 RSVR
LOW

BACK−UP
RSVR LOW

#1 ENG
ANTI−ICE ON

#1 ENG INLET
ANTI−ICE ON

#2 ENG INLET
ANTI−ICE ON

#2 ENG
ANTI−ICE ON

APU GEN ON

PRIME BOOST
PUMP ON

BACK−UP
PUMP ON
#2 TAIL RTR
SERVO ON

TEST /
RESET

AUX FUEL
NO
FLOW

VENT
FAIL

OUTBD
EMPTY

QTY LBS
VENT
OVFL

IMBAL

INBD

EMPTY

NO
FLOW

OUTBD
R

EMPTY

XFER MODE MAN XFER XFER FROM PRESS
AUTO

LEFT

MAN

SEARCH LT
* * ON

LDG LT ON

CARGO
HOOK OPEN

HOOK ARMED

PARKING
BRAKE ON

EXT PWR
CONNECTED

INBD

B
O
T
H

O
F
F

ALL
OUTBD
INBD

RIGHT

OUTBD

OFF

E
12

5
2

1
CRS

1000 FT PER
MIN

OFF

2
3
4

10
9
8

BRT /
DIM

11 12 1

VERTICAL SPEED

DOWN

−4

TOTAL FUEL

TRIM FAIL

LFT PITOT
HEAT

APU ACCUM
LOW

H
D
G

S
HDG

8
6

TEMP
C X 10

PUSH

CHIP INPUT
MDL−LH

COURSE

140

APU ON

VOR
ILS
VOR
ILS

IN. HG

100 FT

#2 ENGINE
OIL PRESS

DC ESS
BUS OFF

30 0

DPLR

1000 FT

PITCH

DPLR

ALT

CODE
OFF

MODE SEL

9 0 1

ROLL

DEG

S
D
T
E
A
G
B
0O
10O
20O
30O
40O

#2 CONV

AC ESS
BUS OFF

120

STAB
POS

#1 CONV

#1 ENGINE
OIL TEMP

140

100

#1 ENGINE
OIL PRESS

ENG

PRESS
PSI X 10

14
GA

TEST

ON
20

25
30

ENGINE
IGNITION

ICS IDENT

NON SECURE RADIOS WILL NOT BE KEYED
WHEN USING ANY SECURE RADIO OR THE
INTERCOM FOR CLASSIFIED COMMUNICATIONS

RADIO FM 1
SW NO. 1

UHF
2

VHF
3

FM2
4

VOR
LOC
AUX

MB
ADF
NAV

GPS

DPLR
GPS

12 27

7
5
9

19
13

1
5

2
21

4
7

22
25

17
18

10
14

AB0823
SA

Figure 2-9. Instrument Panel (Sheet 1 of 4)

2-18

Change 4

TM 1-1520-237-10

RAD ALT

TGT

TEMP
\ X 100

SPEED
\ X 10
11

17
9

9
LWG / m3
10

8
8

% RPM

M
H
L
10 15
T 25 5
20
0
FAIL

PRESS
TO
TEST

7
7

% TRQ

4
2

TGT

B
L
A
D
E

MODE
POWER
ON

D
E
I
C
E

AUTO

TEST
IN

M
T

L
TEST

1 − CHAN − 2

U
AN

140

120

100

150

110

100

105

100

100
40

ABS

ALT

1 4 3

M
PROGRESS

BR
M

NORM
SYNC 1

BLADE DE−ICE TEST
PWR
MAIN
TAIL

0
10

30
40

DEG

S
D
T
E
A
G
B
0O
10O
20O
30O
40O

SET

100

G
S
KIAS
LIMIT
150
100
80
60
45

HDG

NAV

ALT

ROLL

CIS MODE SEL

OAT
RTR

PITCH

3 4

RTR

KM
BACK
CRS

FM
HOME

DPLR

VOR
ILS

BACK
CRS

FM
HOME

VERTICAL SPEED

UP
20

5 4

NORM
ALTR

PLT
CPLT

NORM
ALTR

ADF
VOR

TURN
RATE

CRS
HDG

VERT
GYRO

BRG
2

HDG

DOWN
.5

IN. HG

2 9 9 0

H
D
G

OFF

NA
15

2
3
4

10
9
8

2
3

2
100 FT

COURSE

VOR
ILS

11 12 1

MODE SEL
DPLR

I
R
C
M

ALT

1000 FT

30 0

9 0 1

0
1

SYNC 2
EOT

DI VE

ALT

NAV

SET
PUSH
TO TEST

OFF

0
HDG

FEET

30
NAV

STAB
POS

FT X 100

ATT
CLI MB

KNOTS

OFF
DIGITS

CMD

4
110

120

LOW ROTOR
RPM

130

250
200

130

#2 ENG
OUT
MASTER CAUTION
PRESS TO RESET

FIRE

TEST RTR OVERSPEED

DIMMING

#1 ENG
OUT

RADIO CALL
00 0 00

CRS

PRESS
PSI 5 10

OIL

1000 FT PER
MIN

B
A
DPLR
GPS

GPS

B
L
A
D
E

MODE

1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.

RADAR ALTIMETER
BAROMETRIC ALTIMETER
VERTICAL SPEED INDICATOR
MASTER WARNING PANEL
VERTICAL SITUATION INDICATOR
HORIZONTAL SITUATION INDICATOR
AIRSPEED INDICATOR
STABILATOR POSITION PLACARD
STABILATOR INDICATOR
CIS MODE SELECTOR
VSI / HSI MODE SELECTOR
RADIO CALL PLACARD
PILOT’S DISPLAY UNIT
CLOCK
ICE RATE METER
BLADE DEICE CONTROL PANEL
BLADE DEICE TEST PANEL
INFRARED COUNTERMEASURE CONTROL PANEL
CENTRAL DISPLAY UNIT
RADAR WARNING INDICATOR
AUXILIARY FUEL MANAGEMENT PANEL AFMS
ENGINE IGNITION SWITCH
RADIO SELECT PLACARD
CAUTION / ADVISORY PANEL
SECURE WARNING PLACARD
NVG DIMMING CONTROL PANEL
RAD ALT DIMMING

L
UA

D
E
I
C
E

TEST
IN

POWER
ON
O

AUTO

F
F

M
TEST

NORM
SYNC 1

PROGRESS
BLADE DE−ICE TEST
PWR
MAIN
TAIL

SYNC 2
OAT
RTR

EOT

LWG

PRESS
TO
TEST

RTR

g / m3

M
H
L
10 15
T 25. 5
20
0
FAIL

I
R
C
M

OFF

(ON HELICOPTERS WITH REARRANGED
BLADE DEICE PANELS)

Figure 2-9. Instrument Panel (Sheet 2 of 4)

AB0824
SA

Change 4

2-19

TM 1-1520-237-10

#2 ENG
OUT
MASTER CAUTION
PRESS TO RESET

FIRE

MA WRN CAUT / ADVSY

RAD ALT

RADIO CALL
00 0 00

LOW ROTOR
RPM

FUEL

DIMMING

XMSN

QTY
LB X 100

NVG DIMMING

#1 FUEL LOW

#1 GEN

#2 GEN

#2 FUEL LOW

#1 ENGINE
PRESS

#1 GEN BRG

#2 GEN BRG

#2 ENGINE
PRESS

TEMP
C X 10

% RPM

CMD

CLI M B

200

ABS

% TRQ

TEST RTR OVERSPEED

FT X 100

ATT

130

ALT

100

SET

150
G
S

FEET

SET
PUSH
TO TEST

OFF

30
40
DN

VOR
ILS

BACK
CRS

FM
HOME

30
NA

#1 FUEL
FLTR BYPASS

GUST
LOCK

ANTENNA
EXTENDED

#2 FUEL
FLTR BYPASS

#1 ENGINE
STARTER

#1 OIL
FLTR BYPASS

#2 OIL
FLTR BYPASS

#2 ENGINE
STARTER

110

100

105

100

110

BRG
2

HDG

VERT
GYRO

#1 PRI
SERVO PRESS

#2 HYD
PUMP

#2 PRI
SERVO PRESS

TAIL ROTOR
QUADRANT

IRCM
INDP

AUX FUEL

#1 TAIL
RTR SERVO

MAIN XMSN
OIL TEMP

INT XMSN
OIL TEMP

TAIL XMSN
OIL TEMP

APU OIL
TEMP HI

BOOST SERVO
OFF

STABILATOR

SAS OFF

LFT PITOT
HEAT

FLT PATH
STAB

IFF

RT PITOT
HEAT

CHIP INPUT
MDL−LH

CHIP
INT XMSN

CHIP
TAIL XMSN

CHIP INPUT
MDL−RH
CHIP ACCESS
MDL−RH

CHIP ACCESS
MDL−LH

#1 HYD
PUMP

TRIM FAIL

CHIP MAIN
MDL SUMP

APU
FAIL

MR DE−ICE
FAIL

MR DE−ICE
FAULT

TR DE−ICE
FAIL

ICE
DETECTED

MAIN XMSN
OIL PRESS

#1 RSVR
LOW

#2 RSVR
LOW

BACK−UP
RSVR LOW

#1 ENG
ANTI−ICE ON

#1 ENG INLET
ANTI−ICE ON

#2 ENG INLET
ANTI−ICE ON

#2 ENG
ANTI−ICE ON

APU ON

APU GEN ON

PRIME BOOST
PUMP ON

BACK−UP
PUMP ON

APU ACCUM
LOW

SEARCH LT
ON

LDG LT ON

#2 TAIL RTR
SERVO ON

AIR COND
ON

CABIN HEAT
ON

ANTENNA
RETRACTED

PARKING
BRAKE ON

EXT PWR
CONNECTED

10
8

0
−4

PUSH

0
−4
1

TOTAL FUEL

0
1

1
CRS

TO TEST

0 3
33
0 0 0 0

2
ON

DIM

OFF
DIGITS

BRT /
DIM

2
3
4

10
9
8

6
1000 FT PER
MIN

11 12 1

VERTICAL SPEED

5
2

CRS
HDG

DOWN

TURN
RATE

CHIP
#2 ENGINE

120

ADF
VOR

NORM
ALTR

H
D
G

PLT
CPLT

BATTERY
FAULT

COURSE

NORM
ALTR

BATT LOW
CHARGE

140

5 4

CHIP
#1 ENGINE

30 0

IINS

2
3

3 4

IN. HG

2 9 9 0

#2 ENGINE
OIL TEMP

2
100 FT

#2 ENGINE
OIL PRESS

DC ESS
BUS OFF

ENG
TEMP
C X 10

FM
HOME

IINS

BACK
CRS

0
1000 FT

PITCH

VOR
ILS

ALT

CODE
OFF

MODE SEL

9 0 1

ROLL

DEG

DI VE

NAV
KIAS
LIMIT
150
100
80
60
45

#2 CONV

AC ESS
BUS OFF

120

S
D
T
E
A
G
B
0O
10O
20O
30O
40O

STAB
POS

#1 CONV

#1 ENGINE
OIL TEMP

140

KNOTS

#1 ENGINE
OIL PRESS

27 30
24

250

18 2
1

PRESS
PSI X 10

14
GA

#1 ENG
OUT

OFF

TEST

25
30

CREW
CALL

FLARE

ICS IDENT
RADIO FM 1

UHF

VHF

FM2

SW NO. 1

VOR

LOC

ADF

AUX

NAV

ENGINE
IGNITION

SYSTEMS SELECT
NON SECURE RADIOS WILL NOT BE KEYED
WHEN USING ANY SECURE RADIO OR THE
INTERCOM FOR CLASSIFIED COMMUNICATIONS

12 33
15

7
5
9

16
28

17
25

14
29

30
31

26 24
27

18
19

VG
IINS
ATT

4
7

21
13

DG
IINS
HDG

1
5

8
2

10
14

AA0516_3A
SA

Figure 2-9. Instrument Panel (Sheet 3 of 4)

2-20

TM 1-1520-237-10

POWER
ON

O
F
F

D
E
I
C
E

130

4
0

PWR

NORM
SYNC 1

MAIN

TAIL

OAT

105

RTR

EOT
TGT

130

140

120

110

100

105

100

95
20

9
LWG / m3

PRESS
TO
TEST

M
H
L
10 15
T 25 5
20
0
FA IL

M
A

30
1

100

30
40

HDG

NAV

ALT

HDG

NAV

ALT

CIS MODE SEL

DI VE

9 0 1

ROLL

PITCH

3 4

ALQ 144

ALQ 135

CM CM
JAM INOP

IRCM
INOP

OVER RDR
TEMP INOP

5
25

FM
HOME

NA
15

DPLR

VOR
ILS

BACK
CRS

FM
HOME

IN. HG

2 9 9 0

5 4

H
D
G

VERTICAL SPEED

UP
20

NORM
ALTR

PLT
CPLT

NORM
ALTR

ADF
VOR

TURN
RATE

CRS
HDG

VERT
GYRO

BRG
2

2
3

COURSE

HDG

DOWN
.5
2

TEST
PUSH FOR
STANDBY

BACK
CRS

100 FT

POWER

2
3
4

10
9
8

VOR
ILS

ALQ 156

DPLR
10

T
H
IG
N

ALT

1
2

1000 FT

30 0

KM
60

11 12 1

SET
PUSH
TO TEST

OFF

FEET

30
NAV

MODE SEL

OFF

G
S
KIAS
LIMIT
150
100
80
60
45

ALT

D
A
Y

0
10

DEG

S
D
T
E
A
G
B
0O
10O
20O
30O
40O

ABS

STAB
POS

SET

A 162 G0
S
CW CW
E THRT
JAM

FT X 100

150

I
R
C
M

RETRACT

ATT
CLI M B

ALQ

KNOTS

140

RTR

CMD

200

120

110

SYNC 2

1 − CHAN − 2

BLADE DE−ICE TEST
2

250

% TRQ

TEST RTR OVERSPEED

PROGRESS

% RPM

LOW ROTOR
RPM

TEST

6
6

AUTO

TEST
IN

L
UA

#2 ENG
OUT
MASTER CAUTION
PRESS TO RESET

FIRE

#1 ENG
OUT

DIMMING

9
8
7

RADIO CALL
00 0 00

MODE

RAD ALT

B
L
A
D
E

Ng
SPEED
% X 10
11

TGT
TEMP
C X 100

PRESS
PSI 5 10

OIL

CRS

1000 FT PER
MIN

OFF

EXTEND

OFF

STATUS
FLARE

1. RADAR ALTIMETER
2. BAROMETRIC ALTIMETER
3. VERTICAL SPEED INDICATOR
4. MASTER WARNING PANEL
5. VERTICAL SITUATION INDICATOR
6. HORIZONTAL SITUATION INDICATOR
7. AIRSPEED INDICATOR
8. STABILATOR POSITION PLACARD
9. STABILATOR POSITION INDICATOR
10. CIS MODE SELECTOR
11. VSI / HSI MODE SELECTOR
12. RADIO CALL PLACARD
13. PILOT’S DISPLAY UNIT
14. CLOCK
15. BLADE DEICE CONTROL PANEL
16. BLADE DEICE TEST PANEL
17. ICE RATE METER

18. ALQ−144 INFRARED COUNTERMEASURE
CONTROL PANEL
19. ASE STATUS PANEL
20. ALQ−156 COUNTERMEASURE PANEL
21. CENTRAL DISPLAY UNIT
22. RADAR WARNING INDICATOR
23. ECM ANTENNA SWITCH
24. ENGINE IGNITION SWITCH
25. BEARING DISTANCE HEADING INDICATOR
26. CREW CALL SWITCH / INDICATOR
27. SYSTEM SELECT PANEL
28. CAUTION / ADVISORY PANEL
29. FLARE DISPENSE SWITCH
30. RADIO SELECT PLACARD
31. SECURE RADIO WARNING PLACARD
32. NVG DIMMING CONTROL PANEL
33. RADAR ALTIMETER DIMMING

AA0516_4C
SA

Figure 2-9. Instrument Panel (Sheet 4 of 4)

2-21

TM 1-1520-237-10

stowed position. The windows may be opened to move a
machine gun into position for firing.
2.11.4 Door Locks. Key door locks are installed on
each of the cabin, cockpit and avionics compartment doors.
A common key is used to lock and unlock the doors from
the outside to secure the helicopter. Each crew chief/gunner
sliding window is locked from the inside only.
2.12 CREW SEATS.
2.12.1 Pilots’ Seats.

WARNING
Do not store any items below seats. Seats
stroke downward during a crash and any
obstruction will increase the probability
and severity of injury.
The pilots’ seats provide ballistic protection and can be
adjusted for the pilots’ leg length and height. The pilot’s
seat is on the right side, and the copilot’s is on the left.
Each seat has a one-piece ceramic composite bucket attached to energy absorption tubes. Each seat is positioned
on a track with the bucket directly above a recess in the
cockpit floor. Crash loads are reduced by allowing the seat
and occupant to move vertically as single unit. Occupant
restraint is provided by a shoulder harness, lap belts, and a
crotch belt.

WARNING
To prevent injury to personnel, do not release either the normal or emergency vertical adjust levers unless someone is sitting in the seat. The extension springs are
under load at all times. With seat at lowest position, the vertical preload on the
seat could be as high as 150 pounds. If no
one is in the seat and vertical adjust lever(s) is released, the seat will be snapped
to the highest stop. Anyone leaning over
the seat or with hands on guide tubes
above linear bearings, will be seriously injured.
a. Seat Height Adjustment. Vertical seat adjustment is
controlled by a lever on the right front of the seat bucket.
Springs are installed to counterbalance the weight of the
seat. The lever returns to the locked position when released.
2-22

Change 10

b. Forward and Rear Adjustment. The seat is adjusted
for leg length by a locking lever on the left front of the seat
bucket. The lever is spring-loaded and returns to the locked
position when released.
c. Emergency Tilt Levers. The emergency tilt release
levers are on each side of the seat support frame. The seat
may be tilted back into the cabin for removal or treatment
of a wounded pilot. Seat tilting can be done from the cabin
only when the seat is in the full down and aft position. On
RA-30525 seats, tilting is achieved by pushing the tilt levers in toward center, and then pulling the seat top rearward. On D3801 and D3802 seats, tilting is achieved by
pushing the tilt levers outboard, and then pulling the seat
top rearward.
d. Emergency Vertical Release Lever. The emergency
vertical release lever permits the seat to drop to the lowest
adjustment point for tilting. The emergency vertical release
lever is on the upper center back of RA-30525 seats, and is
actuated by pulling right on the lever. The emergency vertical release pedal is on the lower back of the D3801 and
D3802 seats, and is actuated by pushing down on the foot
operated pedal.
e. Seat Belts. The pilot’s and copilot’s seats each contain a shoulder harness, seat belt, and a crotch strap connected to a common buckle assembly. All belts and straps
have adjustment fittings. The attachment buckle has a
single-point release that will be common in configuration
on the pilot’s and copilot’s seats; they may be of the lift
lever or rotary release configuration, when the lanyard is
pulled or the release is turned all belts and straps will release simultaneously.
2.12.2 Protective Armor. Armor protection is provided
for the body of the pilot and copilot against 7.62 mm rounds
from the side and from the back and below. Armored wings,
attached to the cockpit interior, consist of a sliding panel at
the outboard side of each seat. A release lever at the front
of each panel permits sliding the panel aft to allow rapid
entrance and exit, as well as freedom of movement for the
seat occupant.
2.12.3 Crew Chief/Gunner Seats.

WARNING
Do not store any items below seats. During a crash any obstruction will increase
the probability and severity of injury.

TM 1-1520-237-10

Two outward facing seats (Figure 2-10), one on each
side of the helicopter at the front of the cabin, are for the
crew chief/gunners. Each seat faces a window. Each seat is
a cable-supported steel tube assembly with a fire-resistant,
high strength fabric seat and backrest. Each seat contains
two lower energy attenuators designed and oriented to reduce personal injury in a crash. Each seat has a complete
lap belt and dual torso-restraint shoulder harness attached
to a dual action rotary release buckle. The shoulder harness
is connected to inertia reels on the seat back and bottom.
This gives the wearer freedom to move about his station.
On helicopters equipped with improved crewchief/gunner’s
restraint system, the restraint system is equipped with a
single action rotary release buckle with a guard. A release
plate must be pressed to allow rotation of release, preventing inadvertent handle rotation from contact with equipment, etc. The inertia reel lock control is replaced by a
shorter push/pull manual locking control. Push in and the
inertia reel is manually locked in place. When the control is
pulled out, the reel will lock on sudden pull.
2.13 TROOP PROVISIONS.

WARNING
Do not store any items below seats. During a crash any obstruction will increase
the probability and severity of injury.
UH In addition to crew chief and gunner seats, troop
seats may be installed for 13 persons. Each troop seat has a
belt and shoulder harness for body restraint. The backs and

seat pans are attached through cables to the cabin ceiling
and through cables and rods to seat fittings installed in the
floor. The seats may be installed in any quantity from 1 to
13. Each seat contains two lower energy attenuators designed and oriented to reduce personnel injury. In Row 1,
do not install a passenger/troop seat in the most forward
center position directly behind the cockpit center console.
When seats are removed from the cabin and stowed in the
storage compartment, adjustments must be made for weight
and balance using data in Figures 6-3 and 6-12.
2.13.1 Troop Seat Belt Operation.

1. Extend shoulder strap and attach shoulder strap
fittings to buckle.
2. Extend lap belt and place across body.
3. Place lap belt fitting into buckle and make certain of positive lock.
4. Adjust lap belt tension and shoulder straps for a
comfortable fit.
2.13.2 DF and ECM Operator’s Seats. EH The seats
are similar to the pilot’s and copilot’s seats except that
armored wing protection is not provided.
2.13.3 Observer’s Seat. EH The observer’s seat is
identical to a troop seat (Figure 2-6). It is installed behind,
and to the right, of the DF operator’s seat.

Change 10

2-23

TM 1-1520-237-10

CREW CHIEF / GUNNER’S
SEAT

TROOP
COMMANDER’S
SEAT

TROOP
SEAT
(TYPICAL)
LEFT
GUNNER’S
SEAT

AA0407
SA

Figure 2-10. Troop Seats

2-24

TM 1-1520-237-10

Section II EMERGENCY EQUIPMENT
2.14 FIRE PROTECTION SYSTEMS.
Fire detection and fire extinguishing systems are installed so that a fire may be detected and put out at either
engine or the APU installation, without affecting the remaining two. The engines and APU are monitored by infrared radiation type sensing units, and protected by a main
and reserve high-rate discharge type fire extinguisher installation.
2.14.1 Fire Detection System. A detection system
provides fire warning to the cockpit in case of fire in either
main engine compartment or in the APU compartment. The
system consists of five radiation-sensing flame detectors,
control amplifiers, and a test panel. Two detectors are installed in each main engine compartment and one detector
is in the APU compartment (Figure 2-1). The flame detectors are solid-state photoconductive cells providing continuous volume optical surveillance of the monitored areas. In
case of fire, the detectors react to the infrared radiation and
send a signal to one of the three control amplifiers which in
turn signals the fire warning assembly lighting the proper
T-handle (Figures 2-7 and 2-13). Also, the master FIRE
warning lights will go on if a fire is detected (Figure 2-9).
The detector system automatically resets itself, with warning lights off, when the infrared radiation source ceases to
emit.
2.14.2 Fire Detector Test Panel. A test switch on the
FIRE DETR TEST panel on the upper console (Figure
2-7), when moved to positions 1 or 2, sends a test signal
through the system to put on the fire warning lights and
verify proper system operation to, but not including, the
photo cells. The No. 1 TEST position lights # 1 and # 2
ENG EMER OFF T-handles and APU T-handle and
checks all firewall mounted detectors. The No. 2 TEST
position lights # 1 and # 2 ENG EMER OFF T-handle
only, and checks all deck mounted detectors. The engines
and APU are completely enclosed within their own firewall
compartment, thus reducing the possibility of a false fire
warning from outside sources. Electrical power to operate
the No. 1 and No. 2 detector system is by the dc essential
bus through circuit breakers marked FIRE DET, NO. 1
ENG and NO. 2 ENG, respectively. Power to operate the
APU detector system is by the battery bus through a circuit
breaker marked APU FIRE DET.
2.14.3 Fire Extinguishing Systems. A high-rate discharge extinguishing system provides a two-shot, main and
reserve capability to either main engine compartment or
APU compartment. Two containers are each filled with liq-

uid and charged with gaseous nitrogen. The containers are
mounted above the upper deck, behind the right engine
compartment (Figure 2-1). Both containers have dual outlets, each with its own firing mechanism. Each extinguishing agent container has a pressure gage, easily viewed for
preflight inspection. The system also has a thermal discharge safety port that will cause a visual indicator on the
right side of the fuselage to rupture, indicating that one or
both containers are empty. Electrical power to operate the
No. 1 main and No. 2 reserve outlet valves is by the No. 2
dc primary bus through a circuit breaker, marked FIRE
EXTGH. Power to operate the No. 2 main and No. 1 reserve fire bottles outlet port valves and the directional control valve is by the battery utility bus through a circuit
breaker on the lower console marked FIRE EXTGH.
2.14.4 Fire Extinguisher Arming Levers (THandles). One APU T-handle is on the upper console
(Figure 2-7) marked APU, and two engine fire extinguisher
T-handles are on the engine control quadrant, marked #1
ENG EMER OFF and #2 ENG EMER OFF (Figures 2-4
and 2-13). The handle marked #1 ENG EMER OFF is for
the No. 1 engine compartment, the handle marked #2 ENG
EMER OFF is for the No. 2 engine compartment, and
APU is for the auxiliary power unit compartment. When a
handle is pulled, dc power actuates the fire extinguisher
logic module to select the compartment to which the fire
extinguisher agent is to be directed, and also energizes the
circuit to the fire extinguisher switch. The ends of the
handles house fire detector warning lights.
2.14.5 Fire Extinguisher Control Panel.

WARNING
In case of fire when ac electrical power is
not applied to the helicopter, the reserve
fire extinguisher must be discharged. Fire
extinguisher agent cannot be discharged
into No. 2 engine compartment if ac electrical power is not applied to helicopter.
The switch, marked FIRE EXTGH, on the upper console (Figure 2-7), has marked positions RESERVE-OFFMAIN. The switch is operative only after one of the ENG
EMER OFF or APU lever (T-handle) has been pulled.
When the switch is placed to MAIN, after an ENG EMER
OFF lever has been pulled, the contents of the main fire
extinguisher bottle are discharged into the corresponding

Change 9

2-25

TM 1-1520-237-10

compartment. When the FIRE EXTGH switch is placed to
RESERVE after an ENG EMER OFF lever has been
pulled, the contents of the opposite fire extinguisher bottle
are discharged into the selected compartment. The contents
of the fire extinguisher bottle discharge into the compartment of the last lever pulled.
2.14.6 Crash-Actuated System. A crash-actuated system is part of the fire extinguisher system. An omnidirectional inertia switch is hard-mounted to the airframe to
sense crash forces. Upon impact of a crash of 10 Gs or
more, the switch will automatically fire both fire extinguishing containers into both engine compartments. Electrical
power is supplied from the battery utility bus through a
circuit breaker on the lower console, marked FIRE EXTGH.

a. UH One hand-operated fire extinguisher (Figure 9-1)
is mounted on the cabin wall left of the gunner’s seat. A
second fire extinguisher is on the copilot’s seat.
b. EH Four hand operated portable fire extinguishers are
installed. One mounted on the right gunner window, one on
the left pilot seat, one on the DF operator seat, and one on
the ECM operator seat. The extinguishers are held in place
by a quick-release spring.
2.15 CRASH AXE.
UH One axe (Figure 9-1) is installed between the pilots’
seats in the cabin.

2.16 FIRST AID KITS.

2.14.7 Hand-Operated Fire Extinguishers.

WARNING
Exposure to high concentrations of extinguishing agent or decomposition products
should be avoided. The liquid should not
be allowed to contact the skin; it could
cause frostbite or low temperature burns.

2-26

Change 8

a. UH Three first aid kits (Figure 9-1) are installed, two
on the back of the left pilot seat and one on the back of the
right pilot seat.
b. EH Five first aid kits are installed. One on the back of
the right pilot seat, two on the back of the left pilot seat,
one on the back of the DF operator seat, and one on the
back of the ECM operator seat.

TM 1-1520-237-10

Section III ENGINES AND RELATED SYSTEMS
2.17 ENGINE.
The T700 engine (Figure 2-11), is a front drive, turboshaft engine of modular construction. One is mounted on
the airframe at either side of the main transmission. The
engine is divided into four modules: cold section, hot section, power turbine section, and accessory section.
2.17.1 Cold Section Module. The cold section module
(Figure 2-11), includes the inlet particle separator, the compressor, the output shaft assembly, and line replaceable
units (LRUs). The inlet particle separator removes sand,
dust, and other foreign material from the engine inlet air.
Engine inlet air passes through the swirl vanes, spinning the
air and throwing dirt out by inertial action into the collector
scroll, after which it is sucked through by the engine-driven
blower and discharged overboard around the engine exhaust duct. The compressor has five axial stages and one
centrifugal stage. There are variable inlet guide vanes and
variable stage 1 and stage 2 vanes. LRUs mounted on the
cold section module are the electrical control unit (ECU)
700 , or digital electronic control (DEC) 701C , anti-icing
and start bleed valve, history recorder 700 , or history
counter 701C , ignition system, and electrical cables.
2.17.2 Hot Section Module. The hot section module
(Figure 2-11) consists of three subassemblies; the gas generator turbine, stage 1 nozzle assembly, and combustion
liner. LRUs on the hot section module are ignitors 701C
and primer nozzles and ignitors 700 . The gas generator
turbine consists of a gas generator stator assembly and a
two-stage air cooled turbine rotor assembly which drives
the compressor and the accessory gear box. Stage 1 nozzle
assembly contains air cooled nozzle segments. The nozzle
assemblies direct gas flow to the gas generator turbine. The
combustion liner is a ring type combustor cooled by air
flow from the diffuser case.
2.17.3 Power Turbine Section Module. The power
turbine module (Figure 2-11), includes a two stage power
turbine, exhaust frame, and the shaft and C-sump assembly.
The LRUs mounted on the power turbine section module
are the thermocouple harness, torque and overspeed sensor,
and Np (% RPM 1 or 2) sensor.
2.17.4 Accessory Section Module. The accessory
section module (Figure 2-11) includes the top mounted accessory gear box and a number of LRUs. The LRUs
mounted on the module are the hydromechanical unit
(HMU), engine driven boost pump, oil filter, oil cooler,
alternator, oil and scavenge pump, particle separator

blower, fuel filter assembly, chip detector, oil filter bypass
sensor, radial drive shaft, fuel pressure sensor, and oil pressure sensor.
2.18 ENGINE FUEL SUPPLY SYSTEM.
The engine fuel supply system consists primarily of the
low pressure engine driven boost pump, fuel filter, fuel filter bypass valve, fuel pressure sensor, hydromechanical unit
(HMU), pressurizing and overspeed unit (POU) 700 , or
overspeed and drain valve (ODV) 701C .
2.18.1 Engine Driven Boost Pump. A low pressure
suction engine driven boost pump is installed on the front
face of the engine accessory gear box (Figure 2-11). It assures that the airframe fuel supply system is under negative
pressure, lessening the potential of fire in case of fuel system damage. Lighting of the #1 or #2 FUEL PRESS caution light at idle speed and above could indicate a leak, or
failed engine driven boost pump.
2.18.2 Fuel Filter. The fuel filter is a barrier type full
flow filter with integral bypass. An electrical switch lights
the caution panel #1 FUEL FLTR BYPASS or #2 FUEL
FLTR BYPASS caution light to indicate filter bypass. In
addition, a red button on the filter housing pops out when
filter element differential pressure indicates impending bypass. Power for the fuel filter bypass lights is from the No.
1 and No. 2 dc primary busses through circuit breakers
marked NO. 1 and NO. 2 ENG WARN LTS respectively.
2.18.3 Fuel Pressure Warning System. The engine
fuel pressure warning system for each engine consists of a
pressure switch that turns on the FUEL PRESS caution
light. Fuel pressure caution lights, marked #1 FUEL
PRESS and #2 FUEL PRESS will light when fuel pressure drops below 9 psi. This light can go on when fuel
pressure drops, due to failure of the low pressure boost
pump or an air leak in the suction fuel system. The effect
will vary depending upon the size of the leak. The effect
will be more serious at low engine power. A large enough
leak may cause a flameout. Power for the No. 1 engine fuel
pressure warning system is supplied by the No. 1 dc primary bus through the NO. 1 ENG WARN LTS circuit
breaker. Power for the No. 2 engine fuel pressure warning
system is supplied by the No. 2 dc primary bus through the
NO. 2 ENG WARN LTS circuit breaker.
2.18.4 Engine Fuel System Components. Control of
fuel to the combustion system is done by the HMU. The
HMU, mounted on the rear center of the accessory gear box

Change 9

2-27

TM 1-1520-237-10

OIL
COOLER

FUEL FILTER IMPENDING
BYPASS BUTTON

OIL LEVEL
INDICATOR

BLEED−AIR PORT

FUEL PRESSURE
SENSOR

Np
(%RPM)
SENSOR

INLET PARTICLE
SEPARATOR
BLOWER

ANTI−ICING AND
START BLEED VALVE

MAIN FUEL
NOZZLE

PRIMER FUEL
NOZZLE 700

IGNITER PLUG

LEFT SIDE
ALTERNATOR

OIL FILTER
BYPASS SENSOR

OIL FILTER
BYPASS BUTTON

CHIP DETECTOR

FUEL BOOST
PUMP

OIL TEMPERATURE
SENSOR

FUEL FILTER

SWIRL VANES
OIL PRESSURE
SENSOR

IPS BLOWER
DRAIN LINE
701C
OIL DRAIN PLUG

FRONT VIEW

AA0350_1A
SA

Figure 2-11. Engine T700 (Sheet 1 of 2)
2-28

TM 1-1520-237-10

ACCESSORY SECTION MODULE
THERMOCOUPLE
HARNESS

OIL FILLER
CAP

HYDROMECHANICAL
UNIT

700 HISTORY
RECORDER /
COUNTER 701C

STARTER

IGNITOR
PLUG
TORQUE AND
OVERSPEED SENSOR

HOT SECTION
MODULE (INTERNAL)

IGNITION
EXCITER

700 ECU / DEC 701C

POWER TURBINE MODULE

OIL LEVEL
INDICATOR

COLD SECTION MODULE

RIGHT SIDE

AA0350_2A
SA

Figure 2-11. Engine T700 (Sheet 2 of 2)
(Figure 2-11), contains a high pressure pump that delivers
fuel to the POU/ODV. Various parameters are sensed by
the HMU and influence fuel flow, variable geometry position, and engine anti-ice start bleed valve operation. Fuel
from the HMU flows to a POU 700 or ODV 701C .
2.18.4.1 Pressurizing and Overspeed Unit.
700 The POU sends some of the fuel through the fuel start
manifold tube to the primer nozzles and allows back flow
of high pressure air for purging. The rest of the fuel is sent
through the main fuel manifold to the injectors for starting
acceleration and engine operation. It purges fuel from the
primer nozzles after light off. It purges fuel from the primer

nozzle and main fuel manifold on shutdown. It also reduces
fuel flow to prevent an engine overspeed when the overspeed system is tripped as sensed by the ECU.
2.18.4.2 Overspeed and Drain Valve. 701C The ODV
sends fuel through the main fuel manifold to the injectors
for starting acceleration and engine operations. It purges
fuel from the main fuel manifold and allows back flow of
high pressure air for purging. It shuts off fuel flow to prevent an engine overspeed when the overspeed system is
tripped as sensed by the DEC. It also shuts off fuel to
prevent hot starts when activated by the hot start preventor
(HSP).

2-29

TM 1-1520-237-10

2.19 ENGINE ALTERNATOR.

2.20 IGNITION SYSTEM.

2.19.1 Engine Alternator. 700 The engine alternator
(Figure 2-11) supplies ac power to the ignition exciter and
electrical control unit (ECU). It also supplies a signal to the
Ng SPEED cockpit indicator. All essential engine electrical
functions are powered by the alternator.

The engine ignition system is a noncontinuous ac powered, capacitor discharge, low voltage system. It includes a
dual exciter, two igniter plugs, ignition leads, and ENGINE
IGNITION keylock switch.
2.21 HISTORY RECORDER.

a. When the alternator power supply to the ECU is interrupted, a loss of % RPM 1 or 2 and % TRQ,indications
will occur, with corresponding engine(s) increasing to
maximum power (high side).
b. When the alternator Ng signal is interrupted, a loss of
Ng cockpit indication will occur with a corresponding ENG
OUT warning light and audio.
c. A complete loss of engine alternator power results in
affected engine(s) increasing to maximum power (high
side) with a loss of cockpit indications of % RPM 1 or 2,
% TRQ, and Ng SPEED; and an ENG OUT warning light
and audio will occur. Overspeed protection is still available.

700

The engine history recorder is mounted on the right side
of the swirl frame (Figure 2-11). It displays four digital
counters which records information for maintenance purposes only. The history recorder will only operate with an
ECU.
2.22 HISTORY COUNTER.

701C

The engine history counter is mounted on the right side
of the swirl frame (Figure 2-11). It displays four digital
counters which records information for maintenance purposes only. The history counter will only operate with a
DEC.
2.23 THERMOCOUPLE HARNESS.

2.19.2 Engine Alternator. 701C The engine alternator
(Figure 2-11) supplies ac power to the ignition exciter and
digital electronic control (DEC) unit. It also supplies a signal to the Ng SPEED cockpit indicator. All essential engine electrical functions are powered by the alternator.
a. When the alternator power supply to the DEC is interrupted, 400 Hz 120 VAC aircraft power is utilized to
prevent engine (high side) failure. There will be a loss of
the associated cockpit Ng indication and activation of an
ENG OUT warning light and audio. Overspeed protection
is still available.
b. When the alternator Ng signal is interrupted, a loss of
the associated engine Ng indication, and an ENG OUT
warning light and audio will occur. Because the DEC can
utilize 400 Hz 120 VAC aircraft power, there will be no
loss of associated % RPM 1 or 2 and % TRQ indications.

2-30

Change 8

A seven probe harness measures the temperature of the
gases at the power turbine inlet. It provides a signal to the
ECU 700 , or DEC 701C , that relays it to the history recorder 700 , or history counter 701C , through the signal
data converter (SDC) to the cockpit temperature gage.
2.24 TORQUE AND OVERSPEED AND % RPM
SENSORS.
Two sensors are installed on the exhaust frame of the
engine. One sensor provides the power turbine governor
and tachometer signal to the ECU 700 , or DEC 701C . The
other sensor feeds the torque computation circuit and overspeed protection system.

TM 1-1520-237-10

2.25 ENGINE BLEED-AIR SYSTEM.
Two bleed-air ports are incorporated on the engine. The
outboard port supplies bleed-air to the engine air inlet antiicing system as described in Section III. The inboard port
ties into the pressurized air system. Air from this port is
supplied to the cabin heating system and can be supplied to
the other engine for crossbleed starts.
2.26 ENGINE ANTI-ICING SYSTEMS.
2.26.1 Engine Anti-Icing.

CAUTION

The engine can incur FOD by improper
use of these systems and the other antiice/deice systems. For example, ice shedding off the windshield can cause FOD
damage to the engines.
a. The engine is anti-iced by two systems; the first described in subparagraph b is called an engine anti-ice system and a second described in paragraph 2.26.2 is called
the engine inlet anti-icing. Both of these systems are turned
on by the ENG ANTI-ICE NO. 1 and NO. 2 switches
(Figure 2-7).
b. Engine anti-icing is a combination of bleed-air and
heated engine oil. Anti-icing is controlled by a solenoidoperated air valve. The engine anti-ice/start bleed valve
opens during starting and will remain open at low power
settings until engine reaches 88 to 92% Ng, depending on
the outside air temperature, with anti-ice OFF. The engine
anti-ice/deice system is designed so that in the event of an
electrical failure the valve reverts to the anti-icing mode
and turns on an advisory light indicating #1 ENG ANTIICE ON or #2 ENG ANTI-ICE ON. Axial compressor
discharge air is bled from stage five of the compressor casing, routed through the anti-icing/bleed valve, and delivered to the front frame through ducting. Within the swirl
frame, hot air is ducted around the outer casing to each
swirl vane splitter lip and inlet guide vanes. The hot air is
directed within each vane by a series of baffles. Hot engine
oil passing within the scroll vanes in the main frame prevents ice buildup. Water, snow, and solids are carried out
through the inlet particle separator discharge system.
Switches marked ENG ANTI-ICE NO. 1 or NO. 2 OFF,
and ON, control engine and inlet anti-ice. At the ON position, compressor bleed-air is supplied continuously. Power
to operate the anti-icing system is by the No. 1 and No. 2
dc primary buses respectively, through circuit breakers,
marked NO. 1 ENG ANTI-ICE and ANTI-ICE WARN,

NO. 2 ENG ANTI-ICE and ANTI-ICE WARN respectively.
2.26.2 Engine Inlet Anti-Icing.
a. The engine air inlets are anti-iced by bleed-air from
the engines. Four advisory lights on the caution/advisory
panel, marked #1 ENG ANTI-ICE ON, #2 ENG ANTIICE ON, #1 ENG INLET ANTI-ICE ON and #2 ENG
INLET ANTI-ICE ON are provided for the engines. The
#1 and #2 ENG ANTI-ICE ON advisory lights will go on
when the ENG ANTI-ICE NO. 1 and ENG ANTI-ICE
NO. 2 switches are placed ON. When the anti-ice system is
operating and an engine is started, the inlet anti-ice valve
for that engine will close. The #1 and #2 ENG INLET
ANTI-ICE ON advisory lights operate from temperature
sensed at the engine inlet fairing. When the temperature
reaches about 93°C (199°F), the temperature switch will
turn on the appropriate ENG INLET ANTI-ICE ON advisory light. If this light goes on with the switches at ENG
ANTI-ICE NO. 1 and NO. 2 OFF, it indicates that heat is
being applied to that engine inlet and a malfunction exists.
Inlet anti-icing will turn on if dc primary power failure
occurs; dc electrical power is applied to keep the valve
closed. Functioning of ENG INLET ANTI-ICE is controlled as follows:
(1) Above 13°C (55°F) - Illumination of the ENG INLET ANTI-ICE ON advisory light indicates a system malfunction.
(2) Above 4°C (39°F) to 13°C (55°F) - The ENG INLET ANTI-ICE ON advisory light may illuminate or may
not illuminate.
(3) At 4°C (39°F) and below - Failure of ENG INLET ANTI-ICE ON advisory light to illuminate indicates
a system malfunction. Do not fly the aircraft in known icing
conditions.
b. At engine power levels of 10% TRQ per engine and
below, full inlet anti-ice capability cannot be provided due
to engine bleed limitations. Power to operate the valves is
normally provided from the No. 1 and No. 2 dc primary
buses, respectively, through circuit breakers marked NO. 1
and NO. 2 ENG ANTI-ICE, respectively. During engine
start, power to operate the No. 1 engine inlet anti-ice valve
is provided from the dc essential bus through a circuit
breaker marked NO. 1 ENG START. The #1 and #2 ENG
INLET ANTI-ICE ON advisory lights receive power from
No. 1 and No. 2 dc primary buses, through circuit breakers,
marked NO. 1 and NO. 2 ENG ANTI-ICE WARN, respectively.

Change 10

2-31

TM 1-1520-237-10

2.27 ENGINE OIL SYSTEM.
Lubrication of each engine is by a self-contained, pressurized, recirculating, dry sump system. Included are oil
and scavenge pump, emergency oil system, monitored oil
filter, tank, oil cooler, and seal pressurization and venting.
The oil tank is a part of the main frame. Each scavenge line
has a screen at the scavenge pump to aid fault isolation. A
chip detector with a cockpit warning light is in the line
downstream of the scavenge pump.
2.27.1 Engine Emergency Oil System. The engine
has an emergency oil system in case oil pressure is lost. Oil
reservoirs built into the A and B sumps are kept full during
normal operation by the oil pump. Oil bleeds slowly out of
those reservoirs and is atomized by air jets, providing continuous oil mist lubrication for the bearings. A #1 ENGINE
OIL PRESS or #2 ENGINE OIL PRESS caution panel
light will go on when indicated oil pressure drops below 25
psi on helicopters without modified faceplates on the instrument panel 700 , or below 20 psi on helicopters with modified faceplates 700 , or 22 psi 701C . Power for the caution
lights comes from the No. 1 and No. 2 dc primary buses
through circuit breakers marked NO. 1 and NO. 2 ENG
WARN LTS respectively.
2.27.2 Oil Tank. The oil tank is an integral part of the
engine. Tank capacity is 7 US quarts. The filler port is on
the right. Oil level is indicated by a sight gage on each side
of the tank. Servicing of the tank is required if the oil level
reaches the ADD line. Overservicing is not possible because extra oil will flow out the filler port. The scavenge
pump returns oil from the sumps to the oil tank through six
scavenge screens, each one labeled for fault isolation.
2.27.3 Oil Cooler and Filter. The oil cooler (Figure
2-11) cools scavenge oil before it returns to the tank. Oil
from the chip detector passes through the oil cooler and is
cooled by transferring heat from the oil to fuel. After passing through the oil cooler, oil enters the top of the main
frame where it flows through the scroll vanes. This further
cools the oil and heats the vanes for full-time anti-icing.
The vanes discharge oil into the oil tank. If the oil cooler
pressure becomes too high, a relief valve will open to dump
scavenge oil directly into the oil tank. Oil discharged from
the oil pump is routed to a disposable-element filter. As the
pressure differential across the filter increases, the first indicator will be a popped impending bypass button. As the
pressure increases further, this indication will be followed
by an indication in the cockpit #1 or #2 OIL FLTR BYPASS, after which a filter bypass will occur. Power for the
caution lights is from the No. 1 and No. 2 dc primary buses
respectively, through circuit breakers marked NO. 1 and
NO. 2 ENG WARN LTS. During cold weather starting, or

2-32

on starting with a partially clogged filter, the high-pressure
drop across the filter will cause the bypass valve to open
and the caution lights to go on. The impending bypass indicator has a thermal lockout below 38°C to prevent the
button from popping. A cold-start relief valve downstream
of the filter protects the system by opening and dumping
the extra oil to the gear box case.
2.27.4 Engine Chip Detector. The chip detector is on
the forward side of the accessory gear box. It consists of a
housing with integral magnet and electrical connector, with
a removable screen surrounding the magnet. The detector
attracts magnetic particles at a primary chip detecting gap.
A common oil discharge from the scavenge pump is routed
to a chip detector wired to a cockpit caution light marked
CHIP #1 ENGINE or CHIP #2 ENGINE. If chips are
detected, a signal is sent to the cockpit to light a caution
light, marked CHIP #1 ENGINE or CHIP #2 ENGINE.
Power to operate the engine chip detector system is from
the No. 1 and No. 2 dc primary buses respectively, through
circuit breakers marked WARN LTS, under the general
headings NO. 1 ENG and NO. 2 ENG.
2.28 ENGINE START SYSTEM.
The pneumatic start system uses an air turbine engine
start motor for engine starting. System components consist
of an engine start motor, start control valve, external start
connector, check valves, controls and ducting. Three pneumatic sources may provide air for engine starts: the APU,
engine crossbleed, or a ground source. When the start button is pressed, air from the selected source is directed
through the start control valve to the engine start motor.
The #1 ENGINE STARTER or #2 ENGINE STARTER
caution light will go on at this time and remain on until the
starter drops out. As the engine start motor begins to turn,
an overrun clutch engages causing the engine to motor. As
the engine alternator begins to turn, electrical current is
supplied to the ignition exciter. Ignition will continue until
either the ENGINE IGNITION switch is moved to OFF
or starter dropout occurs. The ENG POWER CONT lever
is advanced to IDLE detent for light-off and acceleration.
A starter speed switch terminates the start cycle when cutoff speed is reached (52% to 65% Ng SPEED) and turns
off the starter caution light and engine ignition. Malfunction of the starter speed switch may be overcome by manually holding the start button pressed until reaching 52% to
65% Ng SPEED. To drop out the starter, manually pull
down on the ENG POWER CONT lever. To abort a start,
pull down on the ENG POWER CONT lever and move to
OFF in one swift movement. Power to operate the No. 1
engine start control valve is from the dc essential bus
through a circuit breaker marked NO. 1 ENG START.
Power to operate the No. 2 engine start control valve is

TM 1-1520-237-10

from the No. 2 dc primary bus through a circuit breaker
marked NO. 2 ENG START CONTR. For the 701C engine only, fuel flow to the engine will be automatically shut
off if TGT TEMP exceeds 900°C during the start
sequence 701C .
2.28.1 Engine Ignition Keylock. An ENGINE IGNITION keylock is installed on the instrument panel (Figure
2-9), to short out and prevent ignition exciter current flow
when the switch is OFF and the starter is engaged. The
switch is marked ENGINE IGNITION OFF and ON.
When the switch is ON, the shorts are removed from both
engine alternators, allowing exciter current to flow when
the engine alternator begins to turn. The ENGINE IGNITION is normally ON during flight and turned OFF at
shutdown. One switch serves both engines. If the switch is
OFF, neither engine can be started, although motoring capability remains. When an engine is to be motored without
a start, make certain the ENGINE IGNITION switch is
OFF. To prevent a possible hot or torching start never turn
the ENGINE IGNITION switch ON after motoring has
started. Abort start procedures must be done to remove excess fuel from the engine if a start was attempted with the
switch OFF.
2.28.2 APU Source Engine Start. The APU provides
an on-board source of air and auxiliary electrical power.
The APU bleed-air output is enough to start each engine
individually at all required combinations of ambient temperatures and enough to start both engines simultaneously
within a reduced range of ambient temperatures (Figure
5-5). The AIR SOURCE HEAT/START switch must be
at APU. Refer to Section XII for complete APU description.
2.28.3 Crossbleed Engine Start System. Crossbleed
engine starts are used when one engine is operating and it is
desired to start the other engine from the bleed-air source of
the operating engine. To make a crossbleed start, the operating engine must be at least 90% Ng SPEED. When the
AIR SOURCE HEAT/START switch is placed to ENG,
both engine crossbleed valves will open. Pressing the start
button for the engine not operating will cause the start valve
for that engine to open at the same time the crossbleed
valve for the starting engine will close, and remain closed
until starter dropout occurs. At 52% to 65% Ng SPEED,
the starting engine start valve will close, stopping bleed-air
flow to the starter. Power to operate the bleed shutoff valve
is from No. 1 dc primary bus through a circuit breaker
marked AIR SOURCE HEAT/START.
2.28.4 External Source Engine Start. The external
start pneumatic port (Figure 2-1) is on the left side of the
fuselage. It is the attachment point for a bleed-air line from

an external source for engine starting or helicopter heating
on the ground. The assembly contains a check valve to
prevent engine or APU bleed-air from being vented. The
external air source pressurizes the start system up to the
engine start control valves, requiring only that electrical
power be applied. If an emergency start is made without ac
electrical power, No. 1 engine must be started first because
the No. 2 engine start control valve will not operate without
dc primary bus power.
2.29 ENGINE CONTROL SYSTEM.
The engine control system consists of the ECU 700
DEC 701C engine quadrant, load demand system and speed
control system.
2.29.1 Electrical Control Unit (ECU). 700 The electrical control unit controls the electrical functions of the engine and transmits operational information to the cockpit. It
is a solid-state device, mounted below the engine compressor casing. The ECU accepts inputs from the alternator,
thermocouple harness, Np (% RPM 1 and 2) sensor, torque
and overspeed sensors, torque signal from opposite engine
for load sharing, feedback signals from the HMU for system stabilization, and a demand speed from the engine
speed trim button. The ECU provides signals to the %
RPM 1 and 2 indicators, % TRQ meter, TGT TEMP
indicator, and history recorder.
NOTE
Phantom torque may be observed on the Pilot Display Unit (PDU) torque display of a
non-operating engine while the aircraft’s
other engine is operating during a ground
run. Phantom torque readings of up to 14%
have been observed on the PDU of the nonoperating engine. During startup of the nonoperating engine, its ECU will produce a
normal, positive torque signal which displays the correct torque signal on the respective PDU.
a. In case of an ECU malfunction, the pilot may override the ECU by momentarily advancing the ENG
POWER CONT lever to the LOCKOUT stop, then retarding it to manually control engine power. To remove the
ECU from lockout, the ENG POWER CONT lever must
be moved to IDLE.
b. The torque matching/load sharing system increases
power on the lower-torque engine to keep engine torques
approximately equal. The system does not allow an engine
to reduce power to match a lower power engine. If an en-

Change 8

2-33

TM 1-1520-237-10

gine fails to the high side, the good engine will only attempt to increase torque upward until its Np is 3% above
the reference Np.
c. The temperature limiting system limits fuel flow
when the requirement is so great that the turbine temperature reaches the limiting value of 837°C to 849°C. Fuel
flow is reduced to hold a constant TGT. It is normal to see
a transient increase above the 850°C TGT TEMP when the
pilot demands maximum power (Figure 5-1 transient limits). TGT limiting does not prevent overtemperature during
engine starts, compressor stall, or when the engine is operated in LOCKOUT (Paragraph 9.3e).

CAUTION

Delay in release of NO. 1 and NO. 2 ENG
OVSP TEST A and TEST B button may
result in Ng recycling below idle, resulting
in subsequent engine stall and TGT increase. To avoid damage, TGT must be
monitored during overspeed check.
d. Overspeed protection protects the power turbine from
destructive overspeeds. The system is set to trigger at
106%61% RPM 1 or 2 and will result in an initial reduced
fuel flow and will cycle until the cause of the overspeed is
removed or % RPM is reduced manually. Two momentary
switches marked NO. 1 and NO. 2 ENG OVSP TEST A
and TEST B on the upper console (Figure 2-7), are used to
check the circuits. Testing individual circuits A and B indicates that those systems are complete and performing correctly. Dual closing of A and B serves to check out the
actual overspeed system itself, the overspeed solenoid and
the POU. This check must be done only on the groundby
designated maintenance personnel. The overspeed protection is not deactivated when in LOCKOUT. Power to operate the overspeed system is from two independent
sources: the engine alternators as the primary source, and
the No. 1 and No. 2 ac primary buses as alternate backup
source in case of alternator failure. Circuit protection is
through circuit breakers marked NO. 1 ENG OVSP and
NO. 2 ENG OVSP.
2.29.2 Digital Electronic Control (DEC).
701C The digital electronic control unit controls the electrical function of the engine and transmits operational information to the cockpit. It contains a microcomputer processor in a conductive composite molded case. The DEC
can be fully powered by either the engine alternator or by
the 400 Hz, 120 VAC aircraft power. It incorporates logic
that will eliminate the torque spike signal during engine
start and shutdown.
2-34

Change 10

a. The DEC accepts inputs from the alternator, thermocouple harness, Np (% RPM 1 and 2) sensor, torque and
overspeed sensors, RPM R sensor and collective position
transducer for improved transient droop response, torque
signal from opposite engine for load sharing, feedback signals from the HMU for system stabilization, and the engine
speed trim button for Np demand speed reference.

CAUTION

Delay in release of TEST A/B button may
result in Ng recycling below idle, resulting
in subsequent engine stall and TGT increase. To avoid damage, TGT must be
monitored during overspeed check.
b. The DEC provides signals to the % RPM 1 and 2
indicators, % TRQ meter, TGT TEMP indicator, and engine history counter. It also provides signal validations or
selected input signals within the electrical control system.
Signals are continuously validated when the engine is operating at idle and above. If a failure occurred on a selected
input signal, the failed component or related circuit will be
identified by a preselected fault code (Figure 2-12) displayed on the engine torque meter. These codes are defined
in terms of engine torque. They are displayed for 4 seconds
on, 2 seconds off, starting with the lowest code and rotating
through all applicable codes, then repeating the cycle. They
will only be displayed 30 seconds after both engines are
shutdown with 400 Hz, 120 VAC power applied. They may
be recalled by maintenance and the engine restarted. The
pilot can suppress the fault code display of an engine by
depressing the associated cockpit overspeed test button
(TEST A/B). The pilot may recall it by again depressing
the associated cockpit overspeed test button.
c. In case of a DEC malfunction, the DEC may be recalled by maintenance only, and the engine restarted once
action has been performed.
d. The torque matching/load sharing system increases
power on the lower-torque engine to keep engine torques
approximately equal. The system does not allow an engine
to reduce power to match a lower power engine. If an engine fails to the high side, the good engine will only attempt to increase torque upward until its Np is 3% above
the reference Np.
e. The transient compensation system provides significant droop improvement during some maneuvers by monitoring engine torque, collective rate of change, and RPM R
speed rate of change.

TM 1-1520-237-10

f. The temperature limiting system limits fuel flow when
the TGT TEMP reaches the dual engine 10 minute limiting value of approximately 866°C. The automatic contingency power limiting will switch to a higher single-engine
2 1/2 minute temperature limiting value of approximately
891°C when the opposite % TRQ is less than 50%. Fuel
flow is regulated to hold a constant TGT. It is normal to see
a transient increase above the 903°C TGT TEMP limit
when the pilot demands maximum power (Figure 5-2 transient limits). TGT limiting does not prevent overtemperature during engine starts, compressor stall, or when the engine is operated in LOCKOUT (Paragraph 9.3e).
g. The hot start prevention system (HSP) is a part of the
DEC. It prevents overtemperature during engine starts. The
HSP system receives Np, Ng, and TGT signals. When Np
and Ng are below their respective hot start reference and

TGT TEMP exceeds 900°C, an output from the HSP system activates a solenoid in the ODV. This shuts off fuel
flow and causes the engine to shut down. The HSP system
requires 400 Hz, 120 VAC aircraft power be provided to
the DEC. The pilot can disable the HSP for emergency
starting purposes by pressing and holding the overspeed
test button (TEST A/B) for the engine being started during
the engine start sequence.
h. Overspeed protection protects the power turbine from
destructive overspeeds. The system is set to trigger at
120%61% RPM 1 or 2 and will result in a fuel flow
shut-off causing the engine to flame out. When % RPM is
reduced below the overspeed limit, fuel flow is returned to
the engine and engine ignition will come on to provide a
relight. This cycle will continue until the overspeed condition is removed. Two momentary switches marked NO. 1

Change 10

2-34.1/(2-34.2 Blank)

TM 1-1520-237-10

% TRQ

1
140

SIGNAL FAILED

DIAGNOSTIC
INDICATIONS
DISPLAYED
AT SHUTDOWN

140

DIAGNOSTIC
INDICATION ON TORQUE METER
( 3%)

DEC

15%

Np DEMAND CHANNEL

25%

LOAD SHARE CHANNEL

35%

TGT CHANNEL

45%

ALTERNATOR POWER

55%

Ng CHANNEL

65%

Np CHANNEL

75%

TORQUE AND OVERSPEED
CHANNEL

85%

HOT START PREVENTION
CHANNEL

95%

AIRCRAFT 400 Hz POWER

105%

COLLECTIVE CHANNEL

115%

125%
AA0517
SA

Figure 2-12. Signal Validation - Fault Codes
and NO. 2 ENG OVSP TEST A and TEST B on the upper
console (Figure 2-7), are used to check the circuits. Testing
individual circuits A and B indicates that those systems are
complete and performing correctly. Dual closing of A and
B switches serve to check out both the overspeed system,
and the overspeed drain valve (ODV). This check must be
done only on the ground by designated maintenance personnel. The overspeed protection is not deactivated when in
LOCKOUT. Power to operate the overspeed system is
from two independent sources: the engine alternators as the
primary source, and the No. 1 and No. 2 ac primary buses
as alternate backup source in case of alternator failure. Circuit protection is through circuit breakers marked NO. 1
ENG OVSP and NO. 2 ENG OVSP.

701C

2.29.3 Engine Control Quadrant. The engine control
quadrant (Figure 2-13) consists of two ENG POWER
CONT levers, two ENG FUEL SYS selector levers, and
two ENG EMER OFF T-handles. A starter button is on
each ENG POWER CONT lever. Each ENG POWER
CONT lever has four positions: OFF-IDLE-FLYLOCKOUT. Movement of the ENG POWER CONT levers moves a cable to mechanically shut off fuel or set
available Ng SPEED. The lever is advanced to FLY for
flight. This ENG POWER CONT lever setting represents
the highest power that could be supplied if demanded.
Power turbine speed (% RPM 1 or 2) is not governed until
the power lever is advanced from IDLE. The engine quadrant secondary stop, two stop blocks, the quadrant assem-

Change 10

2-35

TM 1-1520-237-10

bly, and a latch on each ENG POWER CONT lever prevent moving the levers below IDLE detent. When
shutdown is required, the ENG POWER CONT lever must
be pulled out slightly, at the same time the latch release
must be pressed, then the ENG POWER CONT lever can
be moved below IDLE detent. After being moved momentarily to LOCKOUT, the ENG POWER CONT lever is
used to manually control Ng SPEED and % RPM 1 or 2.
With the ENG POWER CONT lever at LOCKOUT, the
automatic TGT limiting system is deactivated and TGT
must be manually controlled. The overspeed protection system is not deactivated when at LOCKOUT.
2.29.4 Load Demand System. With ENG POWER
CONT lever at FLY, the ECU 700 or DEC 701C and
HMU respond to collective signals to automatically control
engine speed and provide required power. During emergency operations, when the ENG POWER CONT lever is
moved to LOCKOUT and then to some intermediate position, the engine will still respond to collective signals.
2.29.5 Engine Speed Control System. An engine
RPM control switch on the collective grips (Figure 2-14)
controls the speed of both engines simultaneously. There is
no individual trim capability. It is used to supply a signal to
the ECU 700 , or DEC 701C for controlling % RPM 1 and
2 as required. The ENG RPM control switch allows adjustment between 96% and 100%. The pilot can override the
copilot’s control. Power for ENG RPM control system is
from the No. 2 dc primary bus through a circuit breaker
marked SPEED TRIM.
2.30 HOVER INFRARED
SYSTEM (HIRSS).

SUPPRESSOR

SUB-

The hover IR suppressor (Figure 2-2) provides improved
helicopter survivability from heat-seeking missiles throughout the flight envelope. The HIRSS kit has no moving parts.
It contains a three-stage removable core which reduces
metal surface and exhaust gas temperature radiation and
prevents line-of-sight viewing of hot engine surfaces. The
HIRSS channels hot exhaust gasses through the three-stage
core and inner baffle to induce the flow of cooling air from
the engine bay and the inlet scoops. The three-stage core
and inner baffle cold surfaces are coated with lowreflectance material. For further cooling, hot exhaust gas is
ducted outboard and downward by the engine, away from
the helicopter by the exhaust deflector, where additional
cooling air is provided by the main rotor downwash. Installation of each HIRSS module requires removal of the standard engine exhaust module and aft cabin door track fairings. HIRSS modules are installed on the basic airframe
equipped with HIRSS fixed provisions by two airframe
mounts. The aft fairings are installed using existing mount-

2-36

Change 8

ing points and hardware. While operating in a non-hostile
environment, the inner baffle can be removed to enhance
helicopter performance.
2.31 ENGINE INSTRUMENTS.
The instrument displays (Figure 2-9) consist of ENG
OIL TEMP and PRESS, TGT TEMP, gas generator Ng
SPEED, power turbine speed (% RPM 1 or 2), rotor speed
% RPM R, engine torque % TRQ, and FUEL QTY to
provide the pilots with engine and subsystem monitoring.
Continuous indications of those parameters are indicated on
vertical scales, digital readouts and caution lights. Instruments without low range turn-off feature: % TRQ, TGT
TEMP, Ng SPEED, ENG OIL TEMP and XMSN TEMP
will remain on as parameter increases and go out as it decreases (Figure 5-1). Power for lighting the displays is from
the No. 1 and No. 2 ac primary and No. 1 and No. 2 dc
primary buses through the signal data converters.
2.31.1 Engine Oil Temperature Indicator. Each engine has an oil temperature sensor wired through the signal
data converter to a vertical scale instrument, marked ENG
OIL TEMP, on the central display unit; and to an engine
oil temperature caution light, marked ENGINE OIL
TEMP, on the caution/advisory panel.
2.31.2 Engine Oil Pressure Indicator. Each engine
has an engine oil pressure transmitter, downstream of the
oil filter, that sends readings to a vertical scale indicator,
marked ENG OIL PRESS, on the instrument display
panel; and to an engine oil pressure caution light, marked
ENGINE OIL PRESS, on the caution panel. The lower
precautionary and prohibited ranges will go out when
reaching the bottom of the normal range. 700 It may be
possible that during IDLE operations, the ENGINE OIL
PRESS caution light will go on. 700 If ENGINE OIL
PRESS caution light comes on at IDLE, verify oil pressure
is acceptable by setting Ng SPEED at 90%, check that
engine oil pressure is at least 35 psi. As pressure increases
above 100 psi 700 , or 120 psi 701C the respective prohibited scale changes to red.
2.31.3 TGT Temperature Indicator. The TGT indicating system consists of thermocouples transmitting to a TGT
TEMP indicator. The indicator assembly has two digital
readouts that indicate precise temperatures.
2.31.4 Gas Generator Speed (Ng) Indicator. The Ng
speed indicating system shows Ng speed for each engine.
The system consists of one alternator winding and Ng
SPEED vertical scale instrument, on the instrument panel,

TM 1-1520-237-10

SECONDARY
IDLE STOP
FOR POWER LEVER

STARTER
BUTTON

O
F D X
F
I
F
R
D

QUADRANT
COVER

SY
S

O
W

L F
O L
C Y
K
O
U

NO. 1 ENG
FUEL SYS
SELECTOR
LEVER

N
G

#2 ENG
EMER O
FF

F
OF LE
ID

N
T

NO. 2 ENG FUEL
SYS SELECTOR
LEVER

NO. 1 ENG
POWER CONT
LEVER

#1 E

EME

NO. 2 ENG
EMER OFF
T−HANDLE

NO. 1 ENG
EMER OFF
T−HANDLE

NO. 2 ENG
POWER CONT
LEVER

A
CENTER
IDLE
COVER
DETENT
IDLE
STOP
BLOCK

PUSH TO
RELEASE

LEVER ASSY

LATCH

CONT

PULL
DOWN

LOOKING INBOARD
RIGHT SIDE

AA0351A
SA

Figure 2-13. Engine Control Quadrant

2-37

TM 1-1520-237-10

giving percent rpm. Digital readouts for Ng SPEED are at
the lower section of the instrument face plate. The threedigit readouts provide a closer indication of Ng SPEED.
2.31.5 Engine Power Turbine/Rotor Speed Indicator. Power turbine and rotor speed are indicated for each
engine on a single instrument marked % RPM 1 R 2 on the
display panel with three vertical scales (Figure 5-1). Power
turbine speed is indicated in % RPM 1 or 2 and rotor speed
% RPM R. Rotor speed is sensed by a speed sensor on the
right accessory module. Power turbine speed is sensed by a
speed sensor on the engine exhaust frame. At the top of the
panel are three warning lights that indicate varying degrees
of rotor overspeed. These lights remain on, once tripped,
and must be manually reset.

2-38

Change 8

2.31.6 Torque Indicator. The torque indicating system
shows the amount of power the engine is delivering to the
main transmission. A torque sensor mounted on the exhaust
case measures the twist of the power turbine shaft, and
transmits this signal to the ECU 700 , or DEC 701C and
signal data converter into the torque indicator marked %
TRQ on the display panel, displaying readings for both
engines. Digital readouts giving torques for each engine are
at the top of the indicator. A photocell on the lower center
of the display will automatically adjust the lighting of the
% RPM and % TRQ indicators with respect to ambient
light.

TM 1-1520-237-10

Section IV FUEL SYSTEM
2.32 FUEL SUPPLY SYSTEM.
A separate suction fuel system is provided for each engine. Fuel is stored in two interchangeable, crashworthy,
ballistic-resistant tanks. The fuel system consists of lines
from the main fuel tanks, firewall-mounted selector valves,
prime/boost pump and fuel tanks, and engine-driven suction pumps. The prime/boost pump primes all fuel lines if
prime is lost, and also acts as an APU boost for APU starts
and operation. A selector valve, driven by cable from the
ENG FUEL SYS selector lever on the engine control
quadrant (Figure 2-13) permits operation of either engine
from either fuel tank. The engines and APU are suction fed,
the APU is fed from the left main fuel tank by a separate
fuel line. All fuel lines are routed in the most direct manner. The fuel line network includes self-sealing breakaway
valves that contain fuel in case of helicopter crash or malfunction. All engine fuel lines are self-sealing with the exception of the APU fuel line.

gine that normally feeds from the empty or low-level tank
is moved to XFD. This connects that engine to the other
tank through the crossfeed system. A check valve in each
crossfeed line prevents air from an inoperative engine’s fuel
line crossing to the operating one.
2.32.3 Fuel Filter. The engine fuel filter has a bypass
valve and bypass warning device. The filter is mounted on
the forward left side of the engine accessory gear box. An
impending bypass warning is incorporated on the filter
housing in the form of a popout button. The bypass valve
opens to assure continuous fuel flow with a blocked filter.
At the same time the valve opens, an electrical switch
closes to light the #1 or #2 FUEL FLTR BYPASS caution
light. Power to operate the bypass warning system is from
the No. 1 and No. 2 dc primary buses through circuit breakers marked NO. 1 and NO. 2 ENG WARN LTS, respectively.
2.33 ENGINE FUEL PRIME SYSTEM.

2.32.1 Fuel Tanks. Both main fuel tanks are crashworthy, self-sealing and interchangeable. Each tank contains a
pressure refuel/defuel valve, fuel quantity and low-level
sensors, high-level shutoff valve, low-level shutoff valve,
check valve sump drain, and a self-sealing breakaway vent
valve. (Refer to Table 2-4 for tank capacity.) Fuel tank
drains are in the sumps to permit removal of sediment and
water and provide fuel sampling.
2.32.2 Engine Fuel System Selector Control. Each
fuel system has a selector valve which is manually operated
through the ENG FUEL SYS selector lever on the overhead engine control quadrant (Figure 2-13). There is an
ENG EMER OFF T-handle on each side of the quadrant
which is arranged so that pulling the handle engages the
ENG FUEL SYS selector lever, bringing it to OFF. The
ENG FUEL SYS selectors are connected to the fuel selector valves with low-friction flexible push-pull cables. Each
lever can be actuated to three positions: OFF, DIR, and
XFD. With the selectors at OFF, the control valves are
closed, allowing no fuel flow to the engines. When the
selectors are moved forward to DIR, the selector valves are
opened, providing fuel flow for each engine from its individual fuel tank. If a tank is empty, or you wish to equalize
fuel in the tanks, the ENG FUEL SYS selector of the en-

NOTE
Priming engines using sump mounted fuel
boost pumps is described in paragraph
8.41.3.
A toggle switch on the upper console, marked FUEL
PUMP, FUEL PRIME, OFF and APU BOOST (Figure
2-7), when moved to FUEL PRIME, energizes the prime/
boost pump and solenoid valves to each main engine fuel
supply line and to the solenoid valve for the APU fuel feed
system. Advisory panel indication is displayed during this
mode by a light marked PRIME BOOST PUMP ON.
Prime pump capacity is not enough to prime an engine
when the opposite engine is running. Engines should therefore be primed individually with both engines off. The
prime/boost pump is actuated and the engine prime valve is
opened whenever the engine starter is operating. This provides fuel pressure to aid in a successful engine start. When
the engine speed reaches starter dropout speed, engine fuel
prime valve will close and the prime/boost pump will also
stop operating if the FUEL PUMP switch is OFF. Power
to operate the prime boost system is from the battery bus
through a circuit breaker marked FUEL PRIME BOOST.

Change 8

2-39

TM 1-1520-237-10

2.34 FUEL QUANTITY INDICATING SYSTEM.
All internal fuel is continuously gaged with the FUEL
QTY gage system (Figure 2-8). The system consists of two
tank unit sensors (probes), one in each tank, a dual channel
fuel quantity gage conditioner, and a dual channel lowlevel warning system. The tank units are connected to the
fuel quantity gages marked FUEL QTY 1-2 on the central
display panel. A separate total fuel quantity readout numerically displays the total quantity of fuel on board. The
system may be checked out by pressing the FUEL IND
TEST pushbutton on the miscellaneous switch panel. The
vertical scales of the FUEL QTY indicator and the digital
readout should show a change, and the #1 and #2 FUEL
LOW caution lights on the caution/advisory panel should
flash. When the button is released, the scales and digital
readout will return to the original readings. The fuel quantity indicating system is powered by the No. 1 ac primary
bus through a circuit breaker, marked NO. 1 AC INST.
2.34.1 Fuel Low Caution Light. Two low-level sensors, one on each probe, provide signals which activate two
low-level caution lights indicating #1 FUEL LOW or #2
FUEL LOW. Those lights flash when the fuel level decreases to approximately 172 pounds in each tank. The illumination of these lights does not mean a fixed time period
remains before fuel exhaustion, but is an indication that a
low fuel condition exists. The fuel-low caution lights are
powered by the No. 1 dc primary bus through a circuit
breaker marked FUEL LOW WARN.
2.34.2 Fuel Boost Pump. The helicopter fuel system
contains an electrically-operated submerged fuel boost
pump in each fuel tank. When the pumps operate, they
provide pressurized fuel to the engine fuel inlet port. Each

2-40

Change 8

boost pump is controlled by a switch on the FUEL BOOST
PUMP CONTROL panel (Figure 2-8). The two-position
switch for each pump, marked ON-OFF, activates the
pump for continuous operation to maintain a head of fuel
pressure at the engine fuel inlet port, regardless of engine
boost pump discharge pressure. An advisory light near each
control switch indicates pump pressure and operation. A
check valve in each pump discharge line prevents fuel recirculation during fuel boost operation, and prevents loss of
engine fuel line prime. #1 or #2 FUEL PRESS caution
light going on is also an indicator to turn on boost pumps.
Power to operate the boost pumps is provided from the No.
1 and No. 2 ac primary buses through circuit breakers
marked NO. 1 and NO. 2 FUEL BOOST PUMP, respectively.
2.34.3 Refueling/Defueling. A pressure refueling and
defueling system provides complete refueling and defueling
of both tanks from one point on the left side of the helicopter (Figure 2-26). Closed circuit refueling uses the pressure
refueling system and its components. No electrical power is
required for the system during refueling or defueling. The
tank full shutoff valve is float-operated. A dual high-level
shutoff system acts as back up for each other. The two
high-level float valves close, causing a back pressure to the
fueling/defueling valve at the bottom of the tank, closing
the refuel valve. The tank empty automatic shutoff system
is a function of the low-level float valve opening to allow
air to be drawn into the line, closing the defuel valve. A
filler neck between the fuselage contour and the fuel cell is
a frangible (breakaway) connection. Gravity fueling is done
through filler neck on each side of the fuselage for the
respective tanks. Gravity defueling capability is provided
through the drains.

TM 1-1520-237-10

Section V FLIGHT CONTROLS
2.35 FLIGHT CONTROL SYSTEMS.
NOTE
Flight near high power RF emitters such as
microwave antennas or shipboard radar may
cause uncommanded AFCS and/or stabilator
control inputs. Electromagnetic interference
(EMI) testing has shown that the master caution light may illuminate before or simultaneously with any uncommanded stabilator
trailing edge movement, with 4° or 5° of
movement being the maximum.
The primary flight control system consists of the lateral
control subsystem, the longitudinal control subsystem, the
collective pitch control subsystem, and the directional control subsystem. Control inputs are transferred from the
cockpit to the rotor blades by mechanical linkages, and
hydraulic servos. Pilot control is assisted by stability augmentation system (SAS), flight path stabilization (FPS),
boost servos, and pitch, roll and yaw trim. Dual cockpit
controls consist of the cyclic stick, collective stick and pedals. The pilot and copilot controls are routed separately to a
combining linkage for each control axis. Outputs from the
cockpit controls are carried by mechanical linkage through
the pilot-assist servos to the mixing unit. The mixing unit
combines, sums, and couples the cyclic, collective, and yaw
inputs. It provides proportional output signals, through mechanical linkages, to the main and tail rotor controls.
2.35.1 Cyclic Stick. Lateral and longitudinal control of
the helicopter is by movement of the cyclic sticks through
push rods, bellcranks, and servos to the main rotor. Movement in any direction tilts the plane of the main rotor blades
in the same direction, thereby causing the helicopter to go
in that direction. Each cyclic stick grip (Figure 2-14) contains a stick trim switch, marked STICK TRIM FWD, L,
R and AFT, a go around switch, marked GA, trim release
switch, marked TRIM REL, a panel light kill switch,
marked PNL LTS, a cargo release switch, marked
CARGO REL, and a transmitter ICS switch, marked RADIO and ICS. Refer to major systems for a complete description of switches on the cyclic grip.
2.35.2 Collective Pitch Control Stick. The collective
sticks change the pitch of the main rotor blades, causing an
increase or decrease in lift on the entire main rotor disc. A
friction control on the pilot’s lever can be turned to adjust
the amount of friction and prevent the collective stick from
creeping. The copilot’s stick telescopes by twisting the grip
and pushing the stick aft to improve access to his seat. Each

collective stick has a grip (Figure 2-14) with switches and
controls for various helicopter systems. These systems are:
landing light control, marked LDG LT PUSH ON/OFF
EXT and RETR; searchlight controls, marked SRCH LT
ON/OFF , EXT, L, R, and RETR; servo shutoff control
switch, marked SVO OFF 1ST STG and 2ND STG; engine speed trim switch, marked ENG RPM INCR and
DECR; and cargo hook emergency release switch, marked
HOOK EMER REL; HUD control switch, marked BRT,
DIM, MODE, and DCLT. All switches are within easy
reach of the left thumb. For a complete description of
switches and controls, refer to major system description.
2.35.3 Mixing Unit. A mechanical mixing unit provides
control mixing functions which minimizes inherent control
coupling. The four types of mechanical mixing and their
functions are:
a. Collective to Pitch - Compensates for the effects of
changes in rotor downwash on the stabilator caused by collective pitch changes. The mixing unit provides forward
input to the main rotor as collective is increased and aft
input as collective is decreased.
b. Collective to Yaw - Compensates for changes in
torque effect caused by changes in collective position. The
mixing unit increases tail rotor pitch as collective is increased and decreases tail rotor pitch as collective is decreased.
c. Collective to Roll - Compensates for the rolling moments and translating tendency caused by changes in tail
rotor thrust. The mixing unit provides left lateral input to
the main rotor system as collective is increased and right
lateral input as collective is decreased.
d. Yaw to Pitch - Compensates for changes in the vertical thrust component of the canted tail rotor as tail rotor
pitch is changed. The mixing unit provides aft input to the
main rotor system as tail rotor pitch is increased and forward input as tail rotor pitch is decreased.
2.35.4 Collective/Airspeed to Yaw (Electronic Coupling). This mixing is in addition to collective to yaw mechanical mixing. It helps compensate for the torque effect
caused by changes in collective position. It has the ability
to decrease tail rotor pitch as airspeed increases and the tail
rotor and cambered fin become more efficient. As airspeed
decreases, the opposite occurs. The SAS/FPS computer
commands the yaw trim actuator to change tail rotor pitch
as collective position changes. The amount of tail rotor
pitch change is proportional to airspeed. Maximum mixing

2-41

TM 1-1520-237-10

occurs from 0 to 40 knots. As airspeed increases above 40
knots, the amount of mixing decreases until 100 knots, after
which no mixing occurs.

SEARCHLIGHT
SWITCH

HOOK
EMER REL

SRC

H LT

ON
OFF

PUSH
ON
OFF

LDG LT

SERVO
SHUTOFF

BRT

V
1S O O
TS F
TG F
EXT

EXT
RETR

DIM
2N

E
RPNG
M

2.35.6 Tail Rotor Pedals. The pedals contain switches
that, when pressed, disengage the heading hold feature of
FPS below 60 KIAS. Adjustment for pilot leg length is
done by pulling a T-handle, on each side of the instrument
panel, marked PED ADJ. The pedals are spring-loaded and
will move toward the operator when unlocked. Applying
pressure to both pedals simultaneously will move the pedals for desired leg position. The handle is then released to
lock the pedal adjusted position.

RETR
DE

LANDING LIGHT
CONTROL

ENGINE
SPEED
TRIM

SEARCHLIGHT
CONTROL

2.35.5 Tail Rotor Control. The tail rotor control system
determines helicopter heading by controlling pitch of the
tail rotor blades. Inputs by the pilot or copilot to the control
pedals are transmitted through a series of control rods,
bellcranks, a mixing unit, control cables and servos to the
pitch change beam that changes blade pitch angle. Hydraulic power to the tail rotor servo is supplied from No. 1 or
the backup hydraulic systems.

2.36 FLIGHT CONTROL SERVO SYSTEMS.
2.36.1 Primary Servos. Main rotor control loads are reacted by three, two-stage primary servos mounted on the
upper deck above the cabin, forward of the main gear box.
Each primary servo contains two independent, redundant
stages with only the mechanical input linkage in common.
If one stage becomes inoperative due to pressure loss, a
bypass valve within the depressurized stage will open, preventing hydraulic lock. Electrical interlocks prevent both
flight control servos from being turned off simultaneously.
If the input pilot valve to the servo becomes jammed, bypass automatically occurs. Automatic bypass is indicated to
the pilot by lighting of the associated PRI SERVO PRESS
caution light.

COLLECTIVE STICK GRIP
(TYPICAL)

HUD

2.36.2 Tail Rotor Servo. Tail rotor control loads are reacted by a two-stage tail rotor servo mounted on the tail
gear box. With the TAIL SERVO switch at NORMAL,
the first stage of this servo is powered by the No. 1 hydraulic system. When the TAIL SERVO switch is moved to
BACKUP, the second stage is powered by the backup system. Should the first stage become inoperative, the backup
pump will come on and power the second stage. All aerodynamic loads are then reacted by the second stage.

HUD CONTROL
SWITCH

M BRT
O
D
D
C
E
L
T
DIM

(ON HELICOPTERS
MODIFIED BY
MWO 1−1520−237−50−62, HUD)

AA0365_1C
SA

Figure 2-14. Collective and Cyclic Grips
(Sheet 1 of 2)
2-42

Change 9

2.36.3 Flight Control Servo Switch. First and second
stage primary servo systems are controlled by the servo
switch, marked SVO OFF, on the pilot’s and copilot’s collective stick grips (Figure 2-14). The marked switch positions are 1ST STG and 2ND STG. The servo systems nor-

TM 1-1520-237-10

TR
IM
ICK
FWD

L
GA

AFT

O
RG
CA EL.
R

STICK TRIM
GO AROUND
ENABLE SWITCH

CARGO HOOK
RELEASE SWITCH

ICS RADIO
CONTROL
I.C.S.

TRIM
RELEASE
SWITCH

IM
TR EL
R

RADIO

PANEL LIGHTS
KILL SWITCH

PNL
LTS

CYCLIC MOUNTED STABILATOR
SLEW−UP SWITCH

CYCLIC STICK GRIP
(TYPICAL)
AA0365_2
SA

Figure 2-14. Collective and Cyclic Grips (Sheet 2 of 2)
2-43

TM 1-1520-237-10

mally operate with the switch in the unmarked center (on)
position. To turn off the first stage primary servos, the SVO
OFF switch is placed to 1ST STG. To turn off the second
stage servo, the switch is placed to 2ND STG. The systems
are interconnected electrically so that regardless of switch
position, a system will not shut off unless there is at least
2350 psi in the remaining system. The servo shutoff valve
operates on current from the No. 1 and No. 2 dc primary
buses through circuit breakers marked NO. 1 and NO. 2
SERVO CONTR respectively.
2.36.4 Flight Control Servo Low-Pressure Caution
Lights. The first, second, and tail rotor stage servo hydraulic low-pressure caution lights are marked #1 PRI
SERVO PRESS, #2 PRI SERVO PRESS, and #1 TAIL
RTR SERVO, and will go on if the pressure is below its
respective switch setting, or if the servo pilot valve becomes jammed. The servo switches and warning lights operate on direct current from the No. 1 and No. 2 dc primary
buses through circuit breakers, marked NO. 1 SERVO
WARN and NO. 2 SERVO WARN, respectively.
2.36.5 Pilot-Assist Servos. Pilot assist servos are normally powered by the No. 2 hydraulic system. If the No. 2
hydraulic pump fails, the pilot assist servos and pitch trim
actuator are powered by the backup hydraulic pump. The
following units are pilot-assist servos: collective, yaw, and
pitch boost servos, which reduce control forces; and three
(pitch, roll, yaw) SAS actuators which transfer the output
of the SAS controllers into control actuations.
2.36.6 Boost Servo. There are three boost servos, collective, yaw, and pitch, installed between the cockpit controls and mixing unit, which reduce cockpit control forces.
The collective and yaw boost servos are turned on and off
by pressing the button marked BOOST on the AUTO
FLIGHT CONTROL panel (Figure 2-15). The pitch boost
servo is turned on when SAS 1 or SAS 2 is ON. The boost
shutoff valves receive power from the dc essential bus
through a circuit breaker, marked SAS BOOST.
2.36.7 Pilot-Assist Controls. An AUTO FLIGHT
CONTROL panel (Figure 2-15), in the lower console, contains the controls for operating the pilot-assist servos and
actuators. The panel contains SAS 1, SAS 2, TRIM, FPS,
BOOST and the FAILURE ADVISORY/POWER ON
RESET lights/switches. STICK TRIM and TRIM REL
switches on the cyclic sticks, are manually operated by either pilot or copilot.

2-44

Change 9

2.37 AUTOMATIC FLIGHT CONTROL SYSTEM
(AFCS).
The AFCS enhances the stability and handling qualities
of the helicopter. It is comprised of four basic subsystems:
Stabilator, Stability Augmentation System (SAS), Trim
Systems, and Flight Path Stabilization (FPS). The stabilator
system improves flying qualities by positioning the stabilator by means of electromechanical actuators in response to
collective, airspeed, pitch rate and lateral acceleration inputs. The stability augmentation system provides short term
rate damping in the pitch, roll, and yaw axes. Trim/FPS
system provides control positioning and force gradient
functions as well as basic autopilot functions with FPS engaged.
2.37.1 Stability Augmentation System (SAS).
NOTE
As the vertical gyro comes up to speed or
when the system is shutdown, the derived
pitch/roll rate signal which feeds SAS 1 will
cause small oscillations in pitch and roll
SAS actuators. This is a temporary situation
and can be eliminated by turning SAS 1 off.
The SAS enhances dynamic stability in the pitch, roll,
and yaw axes. In addition, both SAS 1 and SAS 2 enhance
turn coordination by deriving commands from lateral accelerometers which together with roll rate signals are sent to
their respective yaw channels automatically at airspeeds
greater than 60 knots. The SAS 1 amplifier circuitry operates on 28 vdc power from the dc essential bus through a
circuit breaker marked SAS BOOST providing excitation
for the electronic components within the amplifier. AC
power from the ac essential bus through a circuit breaker
marked SAS AMPL is also required for normal operation
of the SAS. The SAS amplifier uses the vertical gyro roll
output to derive roll attitude and rate for the roll SAS commands and an ac-powered yaw rate gyro for the yaw SAS
commands. Loss of ac power to the vertical gyro or SAS
amplifier causes erratic operation of SAS 1 due to loss of
the reference for the ac demodulators. When this condition
is encountered, the pilot must manually disengage SAS 1.
In case of a malfunction of the SAS 2 function, the input
will normally be removed from the actuator and the SAS 2
fail advisory light on the AUTOFLIGHT CONTROL
panel will go on. If the malfunction is of an intermittent

TM 1-1520-237-10

nature the indication can be cleared by simultaneously
pressing POWER ON RESET switches. If the malfunction is continuous, the SAS 2 should be turned off. With
SAS 1 or SAS 2 off, the control authority of the stability
augmentation system is reduced by one-half (5% control
authority). Malfunction of the SAS 1 system may be detected by the pilot as an erratic motion in the helicopter
without a corresponding failure advisory indication. If a
malfunction is experienced, SAS 1 should be turned off.
SAS actuator hydraulic pressure is monitored. In case of
loss of actuator pressure, or if both SAS 1 and SAS 2 are
off, the SAS OFF caution light will go on.
2.37.2 Trim System. When the TRIM is engaged on
the AUTO FLIGHT CONTROL panel, the pitch, roll and
yaw trim systems are activated to maintain position of the
cyclic and tail rotor controls. Proper operation of the yaw
trim requires that the BOOST on the AUTO FLIGHT
CONTROL panel be on. The tail rotor and lateral cyclic
forces are developed in the electromechanical yaw and roll
trim actuators. Both yaw and roll trim actuators incorporate
slip clutches to allow pilot and copilot control inputs if
either actuator should jam. The forces required to break
through the clutch are 80-pounds maximum in yaw and 13
pounds maximum in roll. The longitudinal force is developed by an electrohydromechanical actuator operated in
conjunction with the SAS/FPS computer. When the pilot
applies a longitudinal or lateral force to the cyclic stick
with trim engaged, a combination detent and gradient force
is felt. The pilot may remove the force by pressing the
thumb-operated TRIM REL switch on the pilot/copilot cyclic grip. The pedal gradient maintains pedal position
whenever the trim is engaged. By placing feet on the pedals, the pedal switches are depressed and the gradient force
is removed. The pedals may then be moved to the desired
position and released. The pedals will be held at this position by the trim gradient. The pedal trim gradient actuator
also includes a pedal damper. The pedal damper is engaged
continuously, independent of electric power and the TRIM
switch on the AUTO FLIGHT CONTROL panel. Operation of the trim system is continuously monitored by the
SAS/FPS computer. If a malfunction occurs, the SAS/FPS
computer will shut off the trim actuator(s) driving the affected axis, and the TRIM FAIL and FLT PATH STAB
caution light will illuminate. If the malfunction is of an
intermittent nature, the indication may be cleared by simultaneously pressing both POWER ON RESET switches. In
addition to the trim release switch, a four-way trim switch
on each cyclic stick establishes a trim position without releasing trim. With trim engaged, the trim position is moved
in the direction of switch movement. The cyclic is moved
by the trim switch in one direction at a time. When FPS is
engaged, the TRIM switch changes the pitch and roll attitude reference instead of the cyclic stick position reference.

STABILATOR CONTROL
MAN SLEW
UP

TEST

AUTO
CONTROL

O
F
F

ON
DN

R
E
S
E
T

AUTO FLIGHT CONTROL
SAS 1

SAS 2

TRIM

FPS

FAILURE

ADVISORY

CPTR SAS 2

ACCL CLTV

TRIM RGYR

A / S GYRO

BOOST

R
E
S
E
T

POWER ON RESET

AA0366A
SA

Figure 2-15. Automatic Flight Control System
(AFCS) Switch Panel
The trim system release feature permits the pilot or copilot
to fly the helicopter with light stick forces. The push-on/
push-off TRIM switch on the AUTO FLIGHT CONTROL panel or the TRIM REL switches on the pilot/
copilot cyclic grips may be used to release trim. When the
switch is ON, the trim system provides gradient and detent
holding force for pitch, roll, and yaw. When turned off, the
trim system is released and light cyclic control forces are
present.
2.37.3 Flight Path Stabilization (FPS).

Change 10

2-45

TM 1-1520-237-10

a. Proper FPS operation requires that the BOOST,
TRIM and SAS 1 and/or SAS 2 functions have been selected on the AUTO FLIGHT CONTROL panel. Although not required for proper operation, the FPS performance will be improved by the proper operation of the
stabilator in the automatic mode. To use the FPS features,
the pilot first assures that BOOST, SAS and TRIM are on
and operating, and then turns the FPS switch ON. The
desired pitch and roll attitude of the helicopter may be established in one of these ways:
(1) Pressing the STICK TRIM switch to slew the
reference attitude to the desired attitude.
(2) Pressing the TRIM REL switch on the pilot/
copilot cyclic grip, manually flying the helicopter to the desired trim condition, and releasing
the TRIM REL switch.
(3) Overriding the stick trim forces to establish the
desired trim condition, and then neutralizing
stick forces by means of the trim switch.
b. The trim attitude, once established, will be automatically held until changed by the pilot. At airspeeds greater
than 60 knots, the pitch axis seeks to maintain the airspeed
at which the trim is established, by variation of pitch attitude. When pitch attitude is changed by means of the
STICK TRIM switch, there is a delay from the time that
the STICK TRIM switch input is removed until the new
reference airspeed is acquired. This is to allow time for the
helicopter to accelerate or decelerate to the new trim speed.
The yaw axis of the FPS provides heading hold at airspeeds
less than 60 knots and heading hold or turn coordination at
airspeeds greater than 60 knots. For heading hold operation
at airspeeds less than 60 knots, the helicopter is maneuvered to the desired heading with feet on pedals. When
trimmed at the desired heading, the pilot may remove feet
from pedals, at which time the existing heading becomes
the reference, which is automatically held. To change heading, the pilot may activate one or both pedal switches, trim
up on the desired heading and remove feet from pedals. At
airspeeds greater than 60 knots, heading hold will be automatically disengaged, and coordinated turn engaged under
these conditions:
(1) STICK TRIM switch is actuated in the lateral
direction.
(2) TRIM REL switch is pressed and roll attitude
is greater than prescribed limits.
(3) About 1/2 inch cyclic displacement and a roll
attitude of about 1.5°. Heading hold is auto2-46

Change 10

matically reengaged and turn coordination disengaged upon recovery from the turn when the
lateral stick force, roll attitude, and yaw rate are
within prescribed limits.
c. To make a coordinated turn, the pilot enters a turn in
one of these ways:
(1) Changing reference roll attitude by pressing the
STICK TRIM switch in the desired lateral direction.
(2) Pressing TRIM REL switch on the cyclic grip
and establishing the desired bank angle with
feet off pedal switches.
(3) Exerting a lateral force on the cyclic stick to
achieve the desired bank angle, and then neutralizing the force with the STICK TRIM
switch.
(4) Keeping a lateral force on the cyclic stick for
the duration of the turn.
d. In each of these ways the ball should remain automatically centered during the entry and recovery from the
turn. If feet are on the pedals, care must be taken not to
apply too much force to the pedals to oppose their motion.
If the pilot intentionally miscoordinates the helicopter, the
result will be a pedal force roughly proportional to sideslip.
The pilot may release the pedal force by pressing the cyclic
TRIM REL switch with feet on pedals. During transition
through 60 knots airspeed, the pilot may feel a slight pedal
motion due to a switching transient which may occur when
the commanded coordinated turn pedal position differs
slightly from the pilot-commanded position. The FPS monitoring is automatic. If a malfunction is detected, the FLT
PATH STAB caution light will go on and the FPS will
either continue to operate in a degraded mode, such as
without heading hold, or without airspeed hold; or may
cease to function altogether. The pilot must take over
manual flight of the helicopter, and may either turn the FPS
off or evaluate performance to determine the degree and
type of degradation, and continue flight with the remaining
features. To help evaluate the nature of the degradation,
eight failure advisory indicators are displayed on two
FAILURE ADVISORY switches on the flight control
panel. These tell the pilot the type of sensor or actuator
which has experienced the failure. If a light goes on, it may
be turned off by pressing the lighted switch. All failure
advisory lights will be on at initial application of power.
The pilot may attempt to clear the indication of temporary
malfunction by simultaneously pressing both FAILURE
ADVISORY switches. If the FLT PATH STAB caution

TM 1-1520-237-10

light goes off, it may be assumed that normal operation is
restored. All FPS functions are provided by automatically
moving the cockpit controls.
2.38 STABILATOR SYSTEM.
a. The helicopter has a variable angle of incidence stabilator to enhance handling qualities. The automatic mode
of operation positions the stabilator to the best angle of
attack for the existing flight conditions. After the pilot engages the automatic mode, no further pilot action is required for stabilator operation. Two stabilator amplifiers
receive airspeed, collective stick position, pitch rate, and
lateral acceleration information to program the stabilator
through the dual electric actuators. The stabilator is programmed to:
(1) Align stabilator and main rotor downwash in low
speed flight to minimize nose-up attitude resulting from
downwash.
(2) Decrease angle of incidence with increased airspeed to improve static stability.
(3) Provide collective coupling to minimize pitch attitude excursions due to collective inputs from the pilot. Collective position sensors detect pilot collective displacement
and programs the stabilator a corresponding amount to
counteract the pitch changes. The coupling of stabilator position to collective displacement is automatically phased in
beginning at 30 KIAS.
(4) Provide pitch rate feedback to improve dynamic
stability. The rate of pitch attitude change of the helicopter
is sensed by a pitch rate gyro in each of the two stabilator
amplifiers and used to position the stabilator to help dampen
pitch excursions during gusty wind conditions. A sudden
pitch up due to gusts would cause the stabilator to be programmed trailing edge down a small amount to induce a
nose-down pitch to dampen the initial reaction.
(5) Provide sideslip to pitch coupling to reduce susceptibility to gusts. When the helicopter is out of trim in a
slip or skid, pitch excursions are also induced as a result of
the main rotor downwash on the stabilator. Lateral accelerometers sense this out of trim condition and signal the stabilator amplifiers to compensate for the pitch attitude
change (called sideslip to pitch coupling). Nose left (right
slip) results in the trailing edge programming down. Nose
right produces the opposite stabilator reaction.

tuator will restrict total maximum movement of the stabilator to about 35° if failure occurs full down, or about 30°
if failure occurs full up. The stabilator actuators receive
power from the dc essential bus and No. 2 dc primary bus
through circuit breakers marked STAB PWR. Since the dc
essential bus is powered by the battery, it is possible to
manually slew one actuator using battery power only. If the
stabilator is slewed up, regain automatic control by manually slewing stabilator full down, then push AUTO CONTROL RESET twice. Otherwise, when only one actuator
is slewed, it causes a very large mismatch between the two
actuator positions. This is detected by the fault monitor and
shuts down the automatic mode upon attempted engagement. Automatic control function sensors, airspeed sensors,
pitch rate gyros, collective position sensor, and lateral accelerometer receive power from the ac essential bus and
No. 2 ac primary bus through circuit breakers marked
STAB CONTR.
2.38.1 Stabilator Control Panel. The stabilator control
panel (Figure 2-8), on the lower console, provides electrical
control of the stabilator system. The panel contains a MAN
SLEW switch, a TEST button, and AUTO CONTROL
RESET switch with a push-to-reset feature. The automatic
mode will allow the stabilator to be automatically operated
from about 39° trailing edge down to 9° trailing edge up.
Manual operation is also restricted to these limits. If a malfunction occurs in the automatic mode, the system will
switch to manual, ON will go off in the AUTO CONTROL window, and the STABILATOR caution and
MASTER CAUTION lights will go on and a beeping tone
will be heard in the pilot’s and copilot’s headphones. It
may be possible to regain the auto mode by pressing the
AUTO CONTROL RESET. If the automatic mode is regained, ON will appear in the AUTO CONTROL switch
window and the caution lights will go off. The stabilator
automatic mode is held in the energized state within the
stabilator control amplifier. On certain occasions during interruption of dc power, such as switching of generators, it is
possible to have conditions where the stabilator automatic
mode may shut down. If the automatic mode shuts down
during flight because of an ac power failure, the helicopter
shall be slowed to 80 KIAS before power is restored. In this
case the AUTO CONTROL RESET switch may be
pressed to reengage the auto mode. If the automatic mode
is not regained, the MASTER CAUTION must be reset,
which turns off the beeping tone, and the stabilator controlled throughout its range with the MAN SLEW switch.
When initial power is applied to the stabilator system, it
will be in automatic mode. The TEST switch is used to
check the AUTO mode fault detector feature and is inop-

b. The above features are provided via inputs to dual
actuators which position the stabilator. Failure of one ac-

Change 10

2-47

TM 1-1520-237-10

erative above 60 KIAS. When pressed, control of the stabilator should go to the manual mode.
2.38.2 Stabilator Position Indicator. Two STAB
POS indicators (Figure 2-9) are on the instrument panel.
They give pilots a remote indication of stabilator position.
The copilot’s STAB POS indicator may vary from the pilot’s indicator as much as 2°. The indicator range is marked
from 45° DN to 10° up. The stabilator position indicator
system is powered from the ac essential bus 26V through a
circuit breaker marked STAB IND.

2-48

Change 10

2.38.3 Cyclic-Mounted Stabilator Slew Up Switch.
Installed on each cyclic stick below the grip (Figure 2-14)
is a pull-type stabilator manual slew up switch. The switch
provides the pilot and copilot with rapid accessibility to
stabilator slew up. The cyclic slew switch is wired in parallel with the stabilator panel MAN SLEW-UP switch position. When the switch is actuated, the stabilator trailing
edge will begin to move up and continue until the up limit
stop is reached or the switch is released.

TM 1-1520-237-10

Section VI HYDRAULIC AND PNEUMATIC SYSTEM
2.39 HYDRAULIC SYSTEM.
The three hydraulic systems are designed to provide full
flight control pressure. The components of the hydraulic
systems are three hydraulic pump modules, two transfer
modules, a utility module, three dual primary servos, one
dual tail rotor servo, four pilot-assist servos, an APU accumulator, an APU handpump, and a servicing handpump.
There are three hydraulic pressure supply systems, number
1, number 2, and backup. All are completely independent
and each is fully capable of providing essential flight control pressure for maximum system redundancy. Complete
redundancy is accomplished by the backup pump providing
hydraulic power to both number 1 and/or number 2 systems
if one or both pumps fail. If two systems lose pressure,
there will be a slight restriction in the maximum rate of
flight control movement due to only one pump supplying
both stages with hydraulic power. An automatic turnoff feature is provided. When the SVO OFF switch (Figure 2-14)
is moved to 1ST STG or 2ND STG position, that stage of
the primary servos is turned off. When the SVO OFF
switch is moved to 1ST STG, the first stage of the primary
servos is turned off. A malfunction in the second stage will
cause first stage (which was turned off) to automatically
turn back on in case the backup system does not take over
the function of the failed second stage. If the second stage
is initially turned off, the sequence is reversed. An additional hydraulic handpump is provided for APU start system.
NOTE
The following listed caution lights may momentarily flicker when the applicable listed
switch is activated; this is considered normal.

SUBSYSTEM

CAUTION LIGHT

SAS 1 or SAS 2 switch
ON

#2 PRI SERVO PRESS
#2 HYD PUMP
BOOST SERVO OFF

BOOST switch ON

#2 PRI SERVO PRESS
#2 HYD PUMP
SAS OFF

TAIL SERVO switch
BACKUP

#1 PRI SERVO PRESS
#1 HYD PUMP

SUBSYSTEM

CAUTION LIGHT

HYD LEAK TEST
switch NORM after
RESET

#1 and #2 PRI SERVO
PRESS
#1 and #2 HYD PUMP

2.40 HYDRAULIC PUMP MODULES.
The hydraulic pump modules are combination hydraulic
pumps and reservoirs. The No. 1, No. 2, and backup pump
modules are identical and interchangeable with each other.
The No. 1 pump module is mounted on and driven by the
left accessory module of the main transmission. The No. 2
pump module is mounted on and driven by the right accessory transmission module. The backup pump module is
mounted on and driven by an ac electric motor. The reservoir part of each pump module has a level indicator window marked, REFILL, FULL, and EXPANSION. A pressure relief and bleed valve protects the pump from high
pressure in the return system. The pump has two filters: a
pressure filter and a return filter. A red indicator button on
each filter will pop out when pressure goes up 70 610 psi
above normal. The pressure filter has no bypass. The return
filter has a bypass valve that opens when return pressure
reaches 100 610 psi above normal. Each pump has three
check valves: one at the external ground coupling, one at
the pressure side, and one at the return side. A fluid quantity switch, mounted on top of each pump module, senses
fluid loss for that system. When the piston in the pump
module moves down to the REFILL mark, the piston
closes the switch, turning on a caution light marked RSVR
LOW. Each hydraulic pump has two temperature sensitive
labels mounted on the side. When a temperature level is
reached a circle turns black. There are two types of labels
used on the pumps. When the temperature label indicates
that a temperature of 132°C (270°F) has been exceeded, an
entry shall be made on DA Form 2408-13-1. The aircraft
should not be flown until appropriate maintenance action
has been taken.
2.40.1 Number 1 Hydraulic System. Number 1 hydraulic system operates with the rotor turning, and supplies
the first stage of all primary servos and the first stage of the
tail rotor servo. The system components are an integrated
pump module, a transfer module, first stage primary servos,
and first stage tail rotor servo. The primary servos are controlled by the SVO OFF switch (Figure 2-14). The switch
can turn off either first or second stage of the primary servos but not both at the same time. First stage tail rotor

Change 10

2-48.1/(2-48.2 Blank)

TM 1-1520-237-10

servo can be manually turned off by a two-position switch
marked TAIL SERVO, on the miscellaneous switch panel
(Figure 2-8). If the fluid quantity of the number one pump
reservoir becomes low, a microswitch will complete an
electrical circuit to close the first stage tail rotor servo
valve. If fluid continues to be lost and the #1 HYD PUMP
caution light goes on, the first stage tail rotor shutoff valve
will open, allowing backup pressure to supply first stage
tail rotor. The logic modules automatically control the hydraulic system. The tail rotor servo is a two-stage servo but,
unlike the primary servos, only one stage is pressurized at a
time.
2.40.2 Number 2 Hydraulic System. The number 2
hydraulic system, which also operates with the rotor turning, supplies the second stage primary servo and the pilotassist servos. System components are the integrated pump
module, transfer module, second stage primary servos, and
pilot-assist modules. Second stage primary servos can be
manually turned off by the SVO OFF switch. The pilotassist servos cannot be turned off collectively, but SAS,
TRIM and BOOST servos can be manually turned off by
switches on the AUTO FLIGHT CONTROL panel. If
fluid quantity of the number two pump reservoir becomes
low, the pilot-assist servo becomes inoperative. If fluid continues to be lost, the #2 HYD PUMP caution light will go
on.

adequate three-phase ac power source. An internal depressurizing valve in the backup pump module reduces the output pressure of the pump upon startup of the electric motor.
This valve unloads the electric motor by reducing torque
requirement at low rpm. After about 0.5 second when main
generator is operating, or 4 seconds when operating from
APU generator or external power, the valve is closed and
3000 psi pressure is supplied to the hydraulic system. This
sequence reduces the current demand during backup system
startup. Pressure sensing switches in the number 1 and
number 2 transfer modules constantly monitor the pressure
output of the number 1 and number 2 pumps. Loss of pressure initiates the backup operation. The system then provides emergency pressure to maintain full flight control capability. A WOW switch on the left main landing gear
provides automatic operation of the backup pump when the
helicopter is in the air, regardless of BACKUP HYD
PUMP switch position, and disables the backup pump ac
thermal switch. A pressure sensing switch at the tail rotor
monitors supply pressure to the first stage tail rotor servo.
The backup pump can supply pressure to the first stage tail
rotor servo if the number 1 pump loses pressure. This gives
the pilot a backup tail rotor servo even with the loss of the
primary hydraulic supply, or #1 RSVR LOW. If a leak in a
primary servo system depletes the backup system fluid, the
backup reservoir level sensing switch will turn on the
BACK-UP RSVR LOW caution light, and the pilot must
manually turn off the leaking primary system.

2.40.3 Backup Hydraulic System.
2.41 HYDRAULIC LEAK DETECTION/ISOLATION
SYSTEM.
CAUTION

Whenever the No. 1 ac generator is inoperative (failed, or not on line) and the
BACKUP PUMP PWR circuit breaker is
out for any reason, ac electrical power
must be shut off before resetting
BACKUP PUMP PWR circuit breaker.
Otherwise, it is possible to damage the
current limiters.
The backup hydraulic pump system supplies emergency
pressure to the number 1 and/or number 2 hydraulic systems whenever a pressure loss occurs. It also supplies pressure to the number 2 stage of the tail rotor servo in case of
a loss of pressure in the first stage of the tail rotor servo or
#1 RSVR LOW indication. This system supplies hydraulic
pressure to all flight control components during ground
checkout. The backup system also provides a hydraulic
pressure for automatic recharging of the APU start system
accumulator. The backup hydraulic system pump module is
driven by an electric motor which can be powered by any

The leak detection/isolation (LDI) system protects the
flight control hydraulic system by preventing the further
loss of hydraulic fluid in case of a leak. The LDI system
uses pressure switches and fluid level sensors for monitoring pump hydraulic fluid level, and pump pressure for primary and tail rotor servos, and pilot-assist servos. When a
pump module reservoir fluid level switch detects a fluid
loss, the logic module follows the sequence detailed in Figure 2-16 to isolate the leak. To accomplish this, the logic
module operates the required shutoff valve(s) to isolate the
leak and turns on the backup pump when required. In the
cockpit the RSVR LOW caution light for that system
lights. Backup pump and shutoff valve(s) operation is automatic through the logic module. If, after the isolation sequence, the leak continues, the leakage is in the stage 1 or
2 primary servos and the appropriate SVO OFF switch
must be moved to the off position by the pilot. By placing
the HYD LEAK TEST switch to TEST, all leak detection/
isolation system components are checked electrically. After
a leak test has been made, the HYD LEAK TEST switch
must be moved to RESET momentarily, to turn off caution
and advisory lights that were on during the test. The

2-49

TM 1-1520-237-10

LEAKAGE IN NO. 1
HYDRAULIC SYSTEM

PARTIAL LOSS OF
NO.1 RESERVOIR
HYDRAULIC FLUID

ACTUATION OF NO. 1
RESERVOIR LEVEL
SENSING SWITCH

#1 RSVR LOW
CAUTION LIGHT ON

TURNS OFF NO. 1
TAIL ROTOR SERVO

#1 TAIL RTR SERVO
CAUTION LIGHT ON

BACKUP PUMP
TURNED ON

BACK−UP PUMP ON
ADVISORY LIGHT ON

TURNS ON NO. 2
TAIL ROTOR SERVO

#2 TAIL RTR SERVO ON
ADVISORY LIGHT ON

COMPLETE LOSS OF
NO. 1 RESERVOIR
HYDRAULIC FLUID

#1 HYD PUMP
CAUTION LIGHT ON

IF NO OTHER LIGHTS
ON, LEAK IS IN
NO. 1 STAGE TAIL
ROTOR SERVO

BACKUP PUMP SUPPLIES
NO. 1 PRI SERVO AND
NO. 1 TAIL ROTOR SERVO
(NO. 1 TAIL ROTOR
SERVO TURNED BACK ON)
IF NO OTHER LIGHTS ON
LEAKAGE IS UPSTREAM
OF NO. 1 TRANSFER
MODULE

#1 PRI SERVO PRESS
CAUTION LIGHT MAY
MOMENTARILY FLICKER

SEE CHAPTER 5
FOR LIMITATIONS
PARTIAL LOSS OF
BACKUP RESERVOIR
HYDRAULIC FLUID

ACTUATION OF BACKUP
RESERVOIR LEVEL
SENSING SWITCH

LEAKAGE IN 1ST
STAGE PRI SERVO
SEE CHAPTER 9

BACK−UP RSVR LOW
CAUTION LIGHT ON

NO PILOT ACTION

PILOT MOVE SERVO
OFF SWITCH TO
1ST STG

RESULTING CONDITION

1. LOSS OF NO. 1 PRIMARY
SERVO AND NO. 1 AND
NO. 2 TAIL ROTOR
SERVO.
2. CAUTION LIGHTS ON
#1 HYD PUMP,
#1 PRI SERVO PRESS,
#1 TAIL RTR SERVO,
#1 RSVR LOW,
BACK−UP RSVR LOW.
3. NO ADVISORY LIGHTS
ON.

1. LOSS OF NO. 1 PRIMARY
SERVO.
2. CAUTION LIGHTS ON
#1 HYD PUMP,
#1 PRI SERVO PRESS,
#1 RSVR LOW,
BACK−UP RSVR LOW.
3. ADVISORY LIGHT ON
BACK−UP PUMP ON.

AA0404_1A
SA

Figure 2-16. Hydraulic Logic Module Operation Principle (Sheet 1 of 2)

2-50

TM 1-1520-237-10

LEAKAGE IN NO. 2
HYDRAULIC SYSTEM

PARTIAL LOSS OF
NO. 2 RESERVOIR
HYDRAULIC FLUID

ACTUATION OF NO. 2
RESERVOIR LEVEL
SENSING SWITCH

#2 RSVR LOW
CAUTION LIGHT ON
BOOST SERVO OFF, TRIM
FAIL, FLT PATH STAB
SAS OFF
CAUTION LIGHTS ON

TURNS OFF − PILOT
ASSIST SERVOS

IF NO OTHER LIGHTS
ON LEAKAGE IS IN
PILOT−ASSIST AREA

SEE CHAPTER 5 FOR
LIMITATIONS
SEE CHAPTER 9
COMPLETE LOSS OF
NO. 2 RESERVOIR
HYDRAULIC FLUID

BACKUP PUMP
TURNED ON

#2 RSVR LOW
#2 HYD PUMP,
BACK−UP PUMP ON
ADVISORY LIGHTS ON
AND NO OTHER
LEAKAGE IS
UPSTREAM OF NO. 2
TRANSFER MODULE

NO PILOT ACTION

RESULTING CONDITION
1. LOSS OF NO. 2 PRIMARY
SERVOS AND PILOT
ASSIST SERVOS.
2. CAUTION LIGHTS ON
#2 RSVR LOW
#2 HYD PUMP
#2 PRI SERVO PRESS,
BACK−UP RSVR LOW,
SAS OFF,
TRIM FAIL,
BOOST SERVO OFF
FLT PATH STAB.
3. NO ADVISORY LIGHTS
ON.

INCREASED PEDAL
AND COLLECTIVE LOADS

#2 HYD PUMP
CAUTION LIGHT ON

BACK−UP PUMP ON
ADVISORY LIGHT ON

PILOT−ASSIST SERVOS
TURNED ON

#2 PRI SERVO PRESS
CAUTION LIGHT MAY
MOMENTARILY FLICKER

BOOST SERVO OFF / SAS
OFF CAUTION LIGHT OFF

SEE CHAPTER 5
FOR LIMITATIONS
PARTIAL LOSS OF
BACKUP RESERVOIR
HYDRAULIC FLUID

BACK−UP RSVR LOW
CAUTION LIGHT ON

ACTUATION OF LOW−
LEVEL SENSING
SWITCH

LEAKAGE IN NO. 2
PRI SERVO
SEE CHAPTER 9

POWER ON RESET
PRESS TRIM / FLT
PATH STAB
CAUTION LIGHTS OFF

PILOT MOVE
SERVO OFF SWITCH
TO 2ND STG

RESULTING CONDITION
1. LOSS OF NO. 2 PRIMARY
SERVO.
2. PILOT−ASSIST SERVO
PRESSURE SUPPLIED BY
BACK−UP PUMP.
3. CAUTION LIGHTS ON
#2 RSVR LOW,
#2 HYD PUMP,
#2 PRI SERVO PRESS
BACK−UP RSVR LOW.
4. ADVISORY LIGHT ON
BACK−UP PUMP ON.
AA0404_2C
SA

Figure 2-16. Hydraulic Logic Module Operation Principle (Sheet 2 of 2)

2-51

TM 1-1520-237-10

BACK-UP PUMP ON advisory light will remain on for
about 90 seconds. Refer to Chapter 8 Section II for test
procedure. Except for the HYD LEAK TEST switch, the
hydraulic leak system consists of components of 1st stage,
2nd stage and backup hydraulic systems. A WOW switch
contact prevents hydraulic leak tests from being made in
flight. Power to operate the hydraulic leak test system is
from the No. 2 dc primary bus through a circuit breaker,
marked NO. 2 SERVO CONTR and dc essential bus
through a circuit breaker, marked BACKUP HYD
CONTR.
2.42 TRANSFER MODULES.
The No. 1 and No. 2 transfer modules connect hydraulic
pressure from the pump modules to the flight control servos. Each module is an integrated assembly of shutoff
valves, pressure switches, check valves, and restrictors. The
modules are interchangeable.
2.42.1 No. 1 Transfer Module. This module has a
transfer valve, a pressure switch, a 1st stage primary shutoff
valve, a 1st stage tail rotor shutoff valve, a restrictor, and
check valves. The transfer valve is spring-loaded to the
open or normal position. If 1st stage hydraulic pressure is
lost, the valve automatically transfers backup pump pressure to the 1st stage system. The 1st stage primary shutoff
valve lets the pilot or copilot shut off 1st stage pressure to
the primary servos and prevents both stages from being
shut off at the same time. The pressure switch lights the #1
HYD PUMP light on the caution advisory panel when
pressure drops below 2000 psi and also sends a signal to a
logic module that pressure is lost in the 1st stage hydraulic
system. The restrictor allows fluid to circulate for cooling
under no-flow conditions. If a fluid leak develops past the
transfer module, the check valves prevent fluid loss on the
return side of the transfer module.
2.42.2 No. 2 Transfer Module. The No. 2 transfer
module is like the No. 1 module except that it supplies 2nd
stage pressure. The pilot assist shutoff valve turns off pressure to the pilot assist module. The 2nd stage primary servo
shutoff valve turns off pressure to the 2nd stage of the primary servos. The pressure switch turns on the #2 HYD
PUMP caution light on the caution/advisory panel when
2nd stage system pressure is below 2000 psi, and also sends
a signal to a logic module that pressure is lost in the 2nd
stage system.

2-52

Change 6

2.42.3 Utility Module. The utility module connects hydraulic pressure from the backup pump to the No. 1 and
No. 2 transfer modules, the 2nd stage of the tail rotor servo,
and the APU accumulator. A pressure switch on the module
senses the backup pump operating and turns on the
BACK-UP PUMP ON advisory light on the caution/
advisory panel. If the flow rate through the module to the
APU accumulator goes over 1-1/2 gpm, a velocity fuse
shuts off flow.
2.42.4 Logic Modules. Two logic modules, one in the
left relay panel and the other in the right relay panel, are
used to control the operation of the hydraulic systems. The
logic modules continually monitor the operation of the hydraulic systems by inputs received from pressure switches,
fluid level switches on the pump modules, and inputs received from control switches in the hydraulic system. The
outputs of the logic modules will either turn on lights on
the caution/advisory panel notifying the pilot of a failure,
and/or turn off one or more valves due to a system malfunction. All switching functions of the hydraulic logic
modules are automatic, except as shown by a dagger (†)
which indicates crewmember action (Figure 2-16).
2.43 RESERVOIR FILL SYSTEM.
A handpump and manual selector valve are on the right
side upper deck of the helicopter for system servicing. Refer to Figure 2-25 for servicing. The three hydraulic system
reservoir levels can be seen from the fill pump location.
The handpump reservoir contains a sight gage above the
handpump crank. A 1-quart level mark indicates a requirement for refill. Refer to Section XV this chapter for servicing.
2.44 PNEUMATIC SUBSYSTEM.
A pneumatic subsystem operating from bleed-air furnished by the main engines, the APU, or an external pneumatic power source, is used to drive the main engine starter,
for heating system operation and external extended range
tank fuel transfer. Bleed-air from the main engines is used
for engine inlet anti-icing subsystem operation. The heating
subsystem and the extended range fuel tanks use bleed-air
supplied by the main engines during flight, and on the
ground by the main engines, APU, or external source. The
subsystem contains check valves at each bleed-air source,
and a shutoff valve at each main engine.

TM 1-1520-237-10

Section VII POWERTRAIN SYSTEM
2.45 POWERTRAIN.
The powertrain consists of inputs from two engines, a
main transmission, intermediate gear box, tail gear box and
connecting drive shafting. Power from the engines is transmitted to the main transmission module through input modules. The main transmission is mounted on top of the cabin
between the two engines (Figure 2-1). It mounts and powers the main rotor head, changes the angle of drive from the
engines, reduces rpm from the engines, powers the tail rotor drive shaft and drives the accessory module. The main
transmission consists of five modules: two input modules;
the main module; and two accessory modules. The main
transmission has a built-in 3° forward tilt.
2.45.1 Input Module. The input modules are mounted
on the left and right front of the main module and support
the front of the engines. They contain an input bevel pinion
and gear, and a freewheel unit. The freewheel unit allows
engine disengagement during autorotation, or in case of a
nonoperating engine, the accessory module will continue to
be driven by the main rotor. The input module provides the
first gear reduction between engine and main module.
2.45.2 Accessory Module. One accessory module is
mounted on the forward section of each input module. Each
accessory module provides mounting and drive for an electrical generator and a hydraulic pump package. A rotor
speed sensor is mounted on the right accessory module and
provides signals for the VIDS. On the UH-60L an additional rotor speed sensor is mounted on the left accessory
module which provides input signals to the DEC for improved transient droop response.
2.45.3 Main Module. The main module contains the
necessary gearing to drive the main rotor and tail rotor
systems. It provides a reduction in speed from the input
module to the main module and the tail drive shaft.
2.46 MAIN TRANSMISSION LUBRICATION
SYSTEM.

CAUTION

UH−60A EH Prolonged nose-down attitudes
of 5 degrees or more may cause high main
transmission oil temperature.

The transmission incorporates an integral wet sump lubrication system that provides cooled, filtered oil to all
bearing and gears. The ac generators on the accessory modules also receive oil for cooling. Oil under pressure is supplied through internally cored oil lines, except for the pressure and return lines of the oil cooler. Refer to servicing
diagram for oil specification and servicing (Table 2-4). The
lubrication system includes two lubrication pumps that are
combination pressure and scavenge types operating in parallel. The main transmission may run at cruise flight for 30
minutes with loss of all oil. Main transmission oil pressure
may fluctuate when the aircraft is known to be in a nose-up
attitude (i.e., slope landings or hover with an extreme aft
CG). Pressure regulating and bypass valves protect the lube
system by returning excess high pressure oil back to the
inlet side of the pump. A two-stage oil filter and various
strainers in the sump prevent contamination. The oil filter
has a visual impending bypass indicator (red button) that
protrudes when the first stage filter becomes contaminated.
When the button pops the filter element must be replaced to
reset. A thermal lockout prevents button popping when oil
is cold and thick. The oil cooler uses a blower driven by the
tail rotor drive shaft to cool oil before it enters the various
modules. The oil cooler has a thermostatic bypass valve
that directs oil flow around the cooler when the oil temperature is below 71°61°C. Other warning and monitoring systems on the main transmission are: MAIN XMSN OIL
TEMP and PRESS caution lights, and XMSN TEMP and
PRESS oil temperature gages. An oil pressure switch on
the left accessory module, the farthest point from the
pumps, causes the MAIN XMSN OIL PRESS caution
light to go on when the pressure drops to 1462 psi. The
transmission oil temperature warning system is triggered by
an oil temperature switch at the oil cooler input to the main
module, near the tail takeoff drive shaft flange. A caution
light, MAIN XMSN OIL TEMP goes on when transmission oil temperature reaches 120°C. Temperature for the
gage is sensed between the sump and the pump. Pressure
readings are taken at the main module manifold. Electrical
power for the warning systems, except chip detection, is
from the No. 2 dc primary bus, through the MAIN XMSN
circuit breaker on the overhead circuit breaker panel.
2.46.1 Transmission Oil Temperature Indicator.
The transmission oil temperature indicator marked XMSN
TEMP is a part of the central display unit (Figure 2-9).
Refer to Chapter 5 for limitations. Power to operate the
temperature indicator and MAIN XMSN OIL TEMP caution light is provided from the No. 1 and No. 2 ac primary

Change 5

2-53

TM 1-1520-237-10

buses through the signal data converters and the No. 2 dc
primary bus through a circuit breaker, marked MAIN
XMSN.
2.46.2 Transmission Oil Pressure Indicator. The
transmission oil pressure indicator, marked XMSN PRESS,
is a part of the central display unit (Figure 2-9). Refer to
Chapter 5 for limitations. Power to operate the pressure
indicator and MAIN XMSN OIL PRESS caution light is
provided from the No. 1 and No. 2 ac primary buses
through the signal data converter and No. 2 dc primary bus
through a circuit breaker marked MAIN XMSN.
2.46.3 Transmission Chip Detector System. The
transmission chip detector system consists of chip detectors
on the left and right input modules, left and right accessory
modules, the main gear box module, and caution lights
marked CHIP INPUT MDL-LH, CHIP INPUT MDLRH, CHIP ACCESS MDL-LH, CHIP ACCESS
MDL-RH and CHIP MAIN MDL SUMP. These detectors provide warning of chips in any of five areas of the
main transmission system. Each chip detector incorporates
a self-sealing provision so that it can be removed for visual
inspection without loss of oil. The magnetic plugs on each
chip detector attract ferrous particles at any of the detector
locations. The fuzz burn-off feature prevents false warnings
by burning off small chips and fuzz. The fuzz burn-off feature is deactivated when oil temperature reaches 140°C.
Deactivation of the fuzz burn-off feature does not disable
detection and illumination of caution lights. The main
transmission chip detector is also connected to a 30 second
time delay relay to allow small chips and fuzz to burn off
and/or wash away. Chips that are too large to burn off or
wash away trigger the detection system which illuminates a
caution light on the caution/advisory panel. The pilot or
maintenance personnel must check the caution/advisory
panel before removing power to determine the location of
the chip. The system is powered by the dc essential bus
through a circuit breaker on the upper console circuit
breaker panel marked CHIP DET.
2.46.4 Built In Test (BIT) Chip Detectors.
NOTE
The MASTER CAUTION PRESS TO RESET light may or may not extinguish after
being pressed to reset while the chip detectors BIT is in progress.
BIT chip detectors will automatically test for a continuous circuit from the caution/advisory panel to the individual
chip detector when power is first applied. Chip detector

2-54

Change 8

caution lights illuminate during test and extinguish after
successful completion of test. When first placing the BATT
switch ON, the CHIP INPUT MDL-LH, CHIP ACCESS
MDL-LH, CHIP INT XMSN, CHIP TAIL XMSN,
CHIP INPUT MDL-RH, and CHIP ACCESS-RH illuminate immediately for approximately 45-70 seconds and then
extinguish. The CHIP MAIN MDL SUMP caution light
illuminates after a 30 second delay for approximately 30
seconds and then extinguishes. A caution light that does not
illuminate indicates a failed test on its chip detector circuit.
2.47 TAIL DRIVE SYSTEM.
Six sections of drive shaft connect the main module to
the tail rotor gear box. The shafts drive the oil cooler blower
and transmit torque to the tail rotor. Each shaft is dynamically balanced tubular aluminum. Multiple disc (flexible)
couplings between sections eliminate universal joints. The
shafts are ballistically tolerant if hit by a projectile and are
suspended at four points in viscous-damped bearings
mounted in adjustable plates and bolted to fuselage support
brackets.
2.47.1 Intermediate Gear Box. Mounted at the base of
the pylon is the oil-lubricated intermediate gear box (Figure
2-1). It transmits torque and reduces shaft speed from the
main gear box to the tail gear box. The intermediate gear
box may run at cruise flight for 30 minutes, with loss of all
oil. An internal metal fuzz suppression chip/temperature
sensor detects metal particles and gear box overtemperature
conditions, to light caution panel lights marked CHIP INT
XMSN and INT XMSN OIL TEMP.
2.47.2 Tail Gear Box. The oil-lubricated tail gear box
(Figure 2-1) at the top of the tail pylon transmits torque to
the tail rotor head. The gear box mounts the tail rotor,
changes angle of drive and gives a gear reduction. It also
enables pitch changes of the tail rotor blades through the
flight control system. The gear box housing is magnesium.
The tail gear box may run at cruise flight for 30 minutes
with loss of all oil. An internal fuzz suppression metal chip/
temperature sensor detects metal particles, and gear box
overtemperature conditions, to light caution panel lights,
marked CHIP TAIL XMSN and TAIL XMSN OIL
TEMP.
2.47.3 Intermediate and Tail Gear Box Chip/
Temperature Systems. The intermediate and tail gear
boxes contain identical chip/temperature sensors that indi-

TM 1-1520-237-10

cate in the cockpit when the gear box temperature is too
high, or a chip is present. The chip detectors incorporate a
fuzz burn-off feature which eliminates false warning due to
fuzz and small particles. When a chip is detected and will
not burn off, a caution light on the caution/advisory panel
will go on, indicating CHIP INT XMSN or CHIP TAIL
XMSN. Power to operate the chip system is provided from
the dc essential bus through a circuit breaker marked CHIP

DET. The oil temperature sensor is a bimetal strip that
reacts to temperatures. When the oil temperature reaches
140°C a switch closes to turn on a caution capsule in the
cockpit, marked INT XMSN OIL TEMP or TAIL XMSN
OIL TEMP. Power to operate the oil temperature system is
from the No. 2 dc primary bus through a circuit breaker
marked MAIN XMSN.

Change 2

2-54.1/(2-54.2 Blank)

TM 1-1520-237-10

Section VIII MAIN AND TAIL ROTOR GROUPS
2.48 ROTOR SYSTEMS.
The rotor system consists of a main rotor and tail rotor.
Both systems are driven by the engines through the transmission system, with pitch controlled by the flight control
system.
2.49 MAIN ROTOR SYSTEM.
The main rotor system consists of four subsystems: main
rotor blades, hub, flight controls and the bifilar vibration
absorber. Four titanium-spar main rotor blades attach to
spindles which are retained by elastomeric bearings contained in one-piece titanium hub. The elastomeric bearing
permits the blade to flap, lead and lag. Lag motion is controlled by hydraulic dampers and blade pitch is controlled
through adjustable control rods which are moved by the
swashplate. When the rotor is not turning, the blades and
spindles rest on hub mounted droop stops. Upper restraints
called antiflapping stops retain flapping motion caused by
the wind. Both stops engage as the rotor slows down during
engine shutdown. Blade retaining pins can be pulled from
the blade spindle joint and the blades folded along the rear
of the fuselage. The bifilar vibration absorber reduces rotor
vibration at the rotor. The absorber is mounted on top of the
hub and consists of a four arm plate with attached weights.
Main rotor dampers are installed between each of the main
rotor spindles modules and the hub to restrain hunting (lead
and lag motions) of the main rotor blades during rotation
and to absorb rotor head starting loads. Each damper is
supplied with pressurized hydraulic fluid from a reservoir
mounted on the side of each damper. The reservoir has an
indicator that monitors the reserve fluid. When the damper
is fully serviced, the indicator will show full gold.
2.49.1 Main Rotor Blades. Four main rotor blades use
a titanium spar for their main structural member. The structure aft of the spar consists of fiberglass skin, Nomex honeycomb filler and a graphite/fiberglass trailing edge. The
leading edge of each blade has a titanium abrasion strip, the
outboard portion of which is protected by a replaceable
nickel strip. Electro-thermal blankets are bonded into the
blades leading edge for deicing. A Blade Inspection Method
(BIMt) indicator (Figure 2-17), is installed on each blade
at the root end trailing edge to visually indicate when blade

spar structural integrity is degraded. If a spar crack occurs,
or a seal leaks, nitrogen will escape from the spar. When
the pressure drops below minimum the indicator will show
red bands. A manual test lever is installed on each BIMt
indicator to provide a maintenance check. The blades are
attached to the rotor head by two quick-release expandable
pins, that require no tools to either remove or install. To
conserve space, all blades can be folded to the rear and
downward along the tail cone. When mooring, the blades
can be tied down with a fitting on the bottom of each blade.
2.49.2 Main Rotor Gust Lock. The gust lock prevents
the blades from rotating when the helicopter is parked. The
gust lock is designed to withstand torque from one engine
at IDLE, and thus allow engine maintenance checks independent of drive train rotation. The locking system consists
of a locking handle at the rear of the cabin (Figure 2-5), a
GUST LOCK caution light on the caution/advisory panel
(Figure 2-9), and a locking device and teeth on the tail rotor
takeoff flange of the main transmission. The lock shall only
be applied when the rotor system is stationary; it can only
be released when both engines are shut down. Power to
operate the caution light is provided from the No. 1 dc
primary bus through a circuit breaker marked LIGHTS
ADVSY.
2.50 TAIL ROTOR SYSTEM.
A cross-beam tail rotor blade system provides antitorque action and directional control. The blades are of
graphite and fiberglass construction. Blade flap and pitch
change motion is provided by deflection of the flexible
graphite fiber spar. This feature eliminates all bearings and
lubrication. The spar is a continuous member running from
the tip of one blade to the tip of the opposite blade. Electrothermal blankets are bonded into the blade leading edge for
deicing. The tail rotor head and blades are installed on the
right side of the tail pylon, canted 20° upward. In addition
to providing directional control and anti-torque reaction,
the tail rotor provides 2.5% of the total lifting force in a
hover. A spring-loaded feature of the tail rotor control system will provide a setting of the tail rotor blades for balance flight at cruise power setting in case of complete loss
of tail rotor control.

Change 10

2-55

TM 1-1520-237-10

SPAR

MAIN ROTOR
BLADE

BIM
INDICATOR

MANUAL
TEST
LEVER

SERVICE
VALVE

RED

WHITE

YELLOW

NORMAL PRESSURE
(SAFE CONDITION)

SPAR

LOW PRESSURE
(UNSAFE CONDITION)

PRESSURE INDICATOR

AA0518
SA

Figure 2-17. Main Rotor Blade and BIMT System
2.51 TAIL ROTOR QUADRANT/WARNING.
The tail rotor quadrant contains microswitches to turn on
a caution light marked TAIL ROTOR QUADRANT if a
tail rotor cable becomes severed. Spring tension allows the
quadrant to operate in a normal manner. Electrical power to
operate the warning system is provided from No. 1 dc primary bus through a circuit breaker marked T RTR SERVO

2-56

WARN. If the helicopter is shut down and/or hydraulic
power is removed with one tail rotor cable failure, disconnection of the other tail rotor cable will occur when force
from the boost servo cannot react against control cable
quadrant spring tension. The quadrant spring will displace
the cable and boost servo piston enough to unlatch the
quadrant cable.

TM 1-1520-237-10

Section IX UTILITY SYSTEMS
2.52 WINDSHIELD WIPERS.
Two electrically-operated windshield wipers are installed, one on the pilot’s windshield and one on the copilot’s windshield (Figure 2-1). Both wiper arms are driven
by a common motor through flexible drives and converters.
Power to operate the windshield wiper system is from No.
1 ac primary bus through a circuit breaker, marked
WSHLD WIPER.
NOTE
The use of rain repellent on the windshields
will improve visibility above speeds of 50
KIAS. Rain repellent may be locally purchased.
2.52.1 Windshield Wiper Control.

CAUTION

To prevent possible damage to windshield
surface, do not operate windshield wipers
on a dry windshield.
Control of the windshield wipers is through a springloaded rotary switch on the upper console (Figure 2-7). The
switch is labeled WINDSHIELD WIPER, with marked
positions PARK-OFF-LOW-HI. When the switch is
turned from OFF to LOW or HI, the wipers will operate at
the corresponding speed. The wipers will stop at any position when the switch is turned OFF. When the switch is
turned to PARK, the wipers will return to the inboard
windshield frame and stop. When the switch is released, it
will return to OFF.
2.52.2 Windshield Anti-Ice/Defogging System.

CAUTION

Continued use of a faulty windshield antiice system may result in structural damage (delamination and/or cracking) to the
windshield.

Pilot’s, copilot’s and center windshields (on helicopters
equipped with center windshield anti-ice system) are electrically anti-iced and defogged. Transparent conductors imbedded between the laminations provide heat when electrical power is applied. The temperature of each panel is
controlled to a heat level of about 43°C (109°F). The windshield anti-ice system fault monitoring circuit prevents
windshield burnout when the windshield surface heat is
above 43°C (109°F). If heat increases, the monitor circuit
will turn off the system. Three switches, one for the pilot,
one for the copilot and one for the center windshield, (when
equipped) are on the upper console (Figure 2-7) with markings of WINDSHIELD ANTI-ICE PILOT-OFF-ON, and
COPILOT-OFF-ON. On helicopters equipped with center
windshield anti-ice an additional switch to control the center windshield is marked WINDSHIELD ANTI-ICECTR-OFF-ON. Power to operate the anti-icing system is
provided by the No. 1 and No. 2 ac primary buses through
circuit breakers marked PILOT WSHLD ANTI-ICE and
CPLT WSHLD ANTI-ICE. On helicopters equipped with
center windshield anti-ice, pilot and center windshield antiice circuit breakers are marked WINDSHIELD ANTI-ICE
PILOT and CTR. Power to control the anti-ice system is
provided by the No. 1 and No. 2 dc primary buses through
circuit breakers marked CPLT WSHLD ANTI-ICE and
PILOT WSHLD ANTI-ICE respectively. On helicopters
equipped with center windshield anti-ice system, control
circuit breakers for pilot’s and center windshield are on the
No. 2 dc primary bus and are marked WINDSHIELD
ANTI-ICE PILOT and CTR. If the APU generator is the
sole source of ac-generated power, the backup pump and
the windshield anti-ice cannot be used simultaneously.
2.53 PITOT HEATER.
Pitot tube heat is provided by heating elements within
each pitot tube head. Power to operate both heating elements is controlled by a single switch on the upper console,
marked PITOT HEAT OFF and ON. When the switch is
placed ON, current flows to the heating elements. Current
sensors in the circuits sense the current flow and keep the
caution lights, marked LFT PITOT HEAT and RT PITOT HEAT, turned off. If a heating element fails, the
current sensor will detect no current flow, and turn on the
caution light for that pitot tube. Power to operate the pitot
tube heaters is provided from the No. 2 ac primary bus for
the right pitot tube, through a circuit breaker marked RT
PITOT HEAT, and from the No. 1 ac primary bus for the
left pitot tube, through a circuit breaker marked LEFT PI-

Do not allow ice to accumulate on the
windshield, as ice shedding can cause engine FOD.

Change 9

2-57

TM 1-1520-237-10

TOT HEAT. Power to operate the caution lights is provided from the No. 1 dc primary bus through a circuit
breaker, marked NO. 1 ENG ANTI-ICE.
2.54 ROTOR BLADE DEICE KIT.

CAUTION

Blade deice operation with erosion strips
installed may cause blade damage.
The rotor blade deice kit (Figure 2-18) consists of the
following: deice control panel, deice test panel, system controller, power distributor, main and tail sliprings, main and
tail blade heating elements, droop stop heaters, caution
lights, outside air temperature (OAT) sensor, a modified
ambient temperature sense line and an icing rate meter subsystem. The blade deice system provides improved mission
performance in icing conditions by applying controlled
electrical power to integral heating elements in the main
and tail rotor blades, causing the ice bond layer to melt,
allowing symmetrical ice shedding. Droop stop heaters apply heat to the droop stop hinge pins, to prevent icing and
permit proper operation. The heaters are electrically powered continuously whenever the blade deice system is operating, either with the power switch ON, or the system in
the TEST mode. The blade deice system, excluding
element-on-time (EOT) failure, may be ground checked using the APU generator. To prevent generator overload when
only the APU generator is operating, an interlock system is
installed to inhibit blade deice test if the backup pump is
operating. If the backup pump should go on during the test
cycle, the MR DE-ICE FAIL caution light will go on immediately, alerting the crew to an invalid test attempt. The
test cycle must then be initiated again. The OAT sensor,
installed below the windshield, provides a signal to the controller for heating EOT of the rotor blades. The lower the
OAT, the longer EOT will be. To reduce power requirements, the blades are deiced in cycles. Power to operate the
blade deice is provided from the No. 1 and No. 2 ac primary buses and No. 2 dc primary bus through circuit breakers, marked ICE-DET, DE-ICE CNTRLR, and DE-ICE
PWR TAIL ROTOR, on the mission readiness circuit
breaker panel in the cabin. Main blade deice power is
routed through current limiters in the deice junction box.
When one main generator is inoperative, deice power can
be supplied by the APU generator.
2.54.1 Blade Deice System Operation. The ice detector is operational anytime power is applied to the heli-

2-58

Change 8

copter. The ice detector senses ice accumulation on a vibrating probe by a change in probe frequency. The
frequency change is processed by the ice rate meter. The
ice rate meter provides a visual display of icing intensity, T
(trace), L (light) blue, M (moderate) yellow, and H (heavy)
red. Also, the ice rate meter sends a signal to the ICE
DETECTED caution light when the BLADE DE-ICE
POWER switch is off, informing the pilot of the requirement to turn on the system. When the system has been
turned on by placing the POWER switch ON, the ice detector aspirator heater is turned on, and the ICE DETECTED caution light is turned off. If the MODE switch
is at AUTO, the rate meter sends an ice rate signal to the
controller. The controller processes the ice rate signal to
produce heater element-off-time, and the OAT signal to
produce the heater EOT. The controller sends command
signals through the main rotor sliprings to the system distributor which responds to controller signals by switching
power in sequence to the main rotor blade heater zones.
Tail rotor blade power is switched directly by the controller
and sent through the tail rotor sliprings to the tail rotor
blades. A tail blade distributor is not required since the
power is applied to the four tail blades simultaneously. The
deice control panel contains a rotary switch which allows
automatic or manual control of blade heater element-offtime. In AUTO (automatic), the ice rate signal is passed on
to the controller, which results in off-time variations proportional to the ice rate. In MANUAL, T, L, or M, fixed
signals are transmitted to the controller, resulting in fixed
element-off-time. Ice rate subsystem malfunctions are indicated by the appearance of a FAIL flag on the rate meter
face, requiring operation of the blade deice system in one
of the three manual modes. MANUAL mode should also
be used when the rate meter has no indicated malfunction,
but any of these three conditions has occurred: 1. Pilot has
determined by his judgment of ice intensity that the ice rate
system is inaccurate. 2. Torque required has increased to an
unacceptable level. 3. Helicopter vibration has increased to
an unacceptable level. During a single main generator failure, blade deice will be dropped until the APU is started
and the APU generator switch is placed ON. Even though
the APU generator switch is ON and providing power to
the blade deice system, the APU GEN ON advisory light
will not be on because of one main generator operating.
2.54.2 Blade Deice System Control Panel. All controls for operating the rotor blade deice system are on the

TM 1-1520-237-10

TAIL ROTOR BLADE
ELECTROTHERMAL
HEATING ELEMENT
(SAME ON ALL BLADES)

ICE DETECTOR
DROOP STOP
HEATER
(TYPICAL 4)

DISTRIBUTOR
ASSEMBLY

TAIL
SLIPRING
ASSEMBLY

DE−ICE
JUNCTION BOX
OUTSIDE AIR
TEMPERATURE
SENSOR

MAIN
SLIPRING
ASSEMBLY

CONTROLLER

MAIN ROTOR BLADE
ELECTROTHERMAL
HEATING ELEMENT
(SAME ON ALL BLADES)
(ON HELICOPTERS WITH
REARRANGED BLADE DE−ICE
PANELS)

LW C

2
0.

1.0

H
1.5

2.0

POWER
ON

D
E
I
C
E

TEST
IN

M
T

L
UA

FA I L

MODE
AUTO

PRESS
TO
TEST

B
L
A
D
E

g/ m
3

L
M

TEST

ICING RATE METER

PROGRESS

DEICE CONTROL PANEL
BLADE DE−ICE TEST

B
L
A
D
E

AUTO

L
UA

D
E
I
C
E

TEST
IN

M
AN

POWER
ON

L
TEST

PWR

NORM
SYNC 1

MODE

MAIN

TAIL

RTR

SYNC 2
OAT
EOT

M
PROGRESS

DE−ICE TEST PANEL
DEICE CONTROL PANEL
BLADE DE−ICE TEST
NORM
SYNC 1

LW C

PWR
MAIN

TAIL

SYNC 2

PRESS
TO
TEST

L
2
0.

g/ m
3

M
5.

1.0

FA I L

H
1.5

2.0

OAT
EOT

RTR

DE−ICE TEST PANEL

ICING RATE METER

AA0389A
SA

Figure 2-18. Rotor Blade Deice Kit
2-59

TM 1-1520-237-10

BLADE DEICE system control panel (Figure 2-18). Controls are described as follows:
CONTROL/
INDICATOR
POWER switch TEST

FUNCTION
Electrically test main and tail
rotor deice system for one
test cycle.

Turns on power to blade
deice controller and turns off
ICE DETECTED caution
light.

OFF

Turns off deice system.

TEST IN PROGRESS

MODE selector
AUTO
MANUAL
T

Green light goes on during
test cycle. At end of test
cycle, light should go off.
System off-time is controlled
by ice rate signal.
Gives pilot manual control of
system off-time.
Sets a fixed element-off-time
for trace icing.

Sets a fixed element-off-time
for light icing.

Sets a fixed element-off-time
for moderate icing.

2.54.3 Blade Deice Test. The BLADE DE-ICE TEST
panel (Figure 2-18) allows the pilot to check the blade deice system for failures that are otherwise dormant during
the normal TEST mode, but that can allow abnormal operation during use. The panel accomplishes this by introducing selected failure signals into the system and requiring
the deice controller built-in-test circuitry to function in a
specific manner. The blade deice test should be done during
the ground checkout before each flight when blade deice
use is anticipated. In the NORM position, the test panel
allows system test to be done without the introduction of
false failure signals. Thus, the system should complete its
self checkout cycle without failure indications on the caution panel. In the SYNC 1 and SYNC 2 positions, the test
panel interrupts the distributor sync line and provides the

2-60

controller with a false sync input. The controller must interpret these false signals as indications of distributor failure, and produce MR DE-ICE FAIL caution light for both
cases. In the OAT position, the test panel short circuits the
OAT sensor input to the controller. BITE circuitry within
the controller must sense the simulated failure and turn on
both the MR DE-ICE FAIL and TR DE-ICE FAIL caution lights. In the EOT position, the test panel biases BITE
circuitry in the controller and the OAT sensor to simulate
malfunctioning primary EOT timing circuits. The biased
BITE circuit is thus deceived into believing that the primary circuits are in error. The controller must turn on both
the MR DE-ICE FAIL and TR DE-ICE FAIL lights when
this occurs. The test panel also functions automatically during blade deice system use to sense contradictory signals
from the deice power circuits. If electrical power remains
applied to either the main or tail rotor heating elements
after the controller signals a FAIL condition or when the
system is OFF, then the corresponding PWR monitor light
on the BLADE DE-ICE TEST panel turns on. The light
informs the crew that further action is required to isolate
the deice loads indicated. The test panel provides a reliability check of critical deice system functions. The pilot, after
doing the indicated tests properly, can be confident that the
deice system primary and BITE electronics are functioning
within specified tolerances.
2.54.4 Blade Deice Test Panel. The control for selecting test functions of the blade deice system is on the
BLADE DE-ICE TEST panel (Figure 2-18). Two PWR
lights on the panel warn of power malfunctions of the main
and tail rotor deice. Control and indicators are as follows:
CONTROL/
INDICATOR

FUNCTION

NORM

Provides a signal path for
normal operation.

SYNC 1

Provides a signal to the
controller to verify operation
of synchronization check
circuitry when POWER
switch is at TEST.

SYNC 2

Provides an open circuit to
the controller to verify
operation of synchronization
check
circuitry
when
POWER switch is at TEST.

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

OAT

Short circuits the OAT sensor
to check BITE circuit sensing
a fault when POWER switch
is at TEST.

EOT

Disables BITE circuits in
controller and OAT sensor to
simulate a malfunctioning
primary EOT timing circuit
when POWER switch is ON
and MODE select switch is
at M (moderate).

PWR MAIN RTR light

Indicates a malfunction has
occurred in the main rotor
primary power.

PWR TAIL RTR light

Indicates a malfunction has
occurred in the tail rotor primary power.

2.55 BLACKOUT CURTAINS.
Curtains are provided to cover the cabin windows and
the opening between the pilot’s compartment and the cabin.
Velcro tape is bonded to the cabin structure and the curtains
with an adhesive. Loops are attached to the curtains to aid
removal.
2.56 WIRE STRIKE PROTECTION SYSTEM.
On helicopters equipped with wire strike protection provisions, the system (Figure 2-1) is a simple, lightweight,
positive system with no motorized or pyrotechnic components used to cut, break, or deflect wires that may strike the

helicopter in the frontal area between the tires and fuselage,
and between the fuselage and main rotor in level flight. The
system consists of nine cutters/deflectors located on the fuselage and landing gear/support. They are: upper cutter on
the rear of the sliding fairing, the pitot cutter/deflector on
the front of the sliding fairing, windshield post and wiper
deflectors, door hinge deflector, step extension and step deflector, landing gear joint deflector, main landing gear
cutter/deflector, and tail landing gear deflector.
2.57 FLIGHT DATA RECORDER (ON HELICOPTERS EQUIPPED WITH FLIGHT DATA RECORDER
KIT).
The flight data recorder system installed in the aft transition avionics compartment is a crash survivable digital
tape recorder providing 25 hours of recorded data on a
continuous loop magnetic tape. Flight data input to the recorder is sent from different locations throughout the helicopter. The recorder begins to record data as soon as ac and
dc essential power is supplied to the helicopter. Electrical
power to operate the data recorder system is provided from
the dc essential bus and ac essential bus through circuit
breakers marked FLT REC on the mission readiness circuit breaker panel. There are no controls provided to the
pilot or copilot for control of the recorder.
2.58 DATA COMPARTMENTS.
Data Compartments are on each cockpit door (Figure
2-4).
2.58A SNOW SKIS.
The skis for the UH-60A/L are designed to keep the
aircraft from becoming immobile when operating on snow
(winter) and tundra (summer).

Change 9

2-61

TM 1-1520-237-10

Section X HEATING, VENTILATING, COOLING,
AND ENVIRONMENTAL CONTROL UNIT
2.59 HEATING SYSTEM.
The subsystem consists of a heated air source, cold air
source, mixing unit, temperature sensing unit, overtemperature sensor, controls, ducting and registers. The heating
system is a bleed-air system and bleed-air supplied in flight
by the main engines, and on the ground by the main engines or the APU. An external connector allows connection
of an external ground source in to the pneumatic system,
that can provide heat when connected. Power to operate
electrical components of the heating system is by the No. 1
dc primary bus through a circuit breaker, marked AIR
SOURCE HEAT/START.
2.59.1 Winterized Heater. The winterized heater consists of a high bleed-air flow mixing valve and a modulation valve. The mixing valve is of enough capacity to keep
the interior temperature of the helicopter at 4°C (39°F), to
ambient temperatures down to -54°C (-65°F). The mixture
sensor controls air mixing to allow control of temperature
used for cabin heat. Bleed-air is mixed with ambient air to
get the desired temperature selected by the variable temperature control on the HEATER control panel (Figure
2-7). Bleed-air is regulated with the modulation valve for
downstream mixing with ambient air when the HEATER
control switch is ON. Overtemperature is prevented by two
overtemperature sensors that deenergize solenoid valves
when bleed-air temperature reaches about 90° to 96°C
(194° to 205°F) at the inlet to the mixing valve or in the
mixing chamber. The temperature sensors control current
flow to the on-off solenoid and the winterization solenoid to
hold them energized, allowing bleed-air to flow to the mixing chamber. When the ENG ANTI-ICE switch is placed
ON or a dc power failure occurs, the winterization solenoid
will deenergize. An interlock system between engine antiice system and the heater winterization solenoid valve prevents engine overbleed by reducing bleed-air flow to the
heater when an ENG ANTI-ICE switch is ON. Operation
of the winterization heating system is the same as in Paragraph 2.59.3.
2.59.2 Heat and Ventilation Controls. A variable
control air mixing valve assembly is used to control the
temperature of air for cabin heating in the helicopter.
Bleed-air from the engine, APU, or external source is mixed
with ambient air to obtain the desired temperature determined by the setting of the sensor in the downstream air
flow. Regulation of the diaphragm position is by a solenoid.
Should the HEATER control switch (Figure 2-7) be turned
OFF, or dc power fail, bleed-air will shut off. The valve

2-62

also has a thermal protective switch that deenergizes the
solenoid if mixed air temperature is over 90° to 96°C (194°
to 205°F). The mixture temperature sensor downstream of
the mixing valve regulates flow output temperature. The
sensor is regulated from the cockpit through a control linkage at the overhead console. The temperature control is
marked HEATER OFF, MED, and HI. Ventilation is controlled through a panel on the upper console marked VENT
BLOWER. When the switch is placed ON, dc power to the
solenoid allows bleed-air to mix with outside air.
2.59.3 Normal Operation.
1. APU or engine - Start (Refer to paragraph 8.22
or 8.23).
2. AIR SOURCE HEAT/START switch - As required. ENG if engine is operating; OFF for
heat from external air source.
3. HEATER ON-OFF switch - ON.
4. VENT BLOWER switch - OFF for maximum
heat.
5. HEATER control - As desired.
2.60 VENTILATION SYSTEM.
2.60.1 Ventilation System. UH The helicopter is ventilated by an electrically-operated blower system controlled
through the VENT BLOWER control panel on the upper
console (Figure 2-7). The VENT BLOWER switch is
marked OFF and ON. When ON, the blower forces ambient air into the cabin ducts. The No. 2 ac primary bus
powers the blower through a circuit breaker, marked HEAT
& VENT. It is also controlled by dc power from the No. 2
dc primary bus through the VENT BLOWER switch protected by a circuit breaker, marked HEAT VENT. Ram air
vents for cooling the cockpit area are on each side of the
upper console and at the front of the lower console (Figure
2-4) and are controlled by turning the nozzle to control the
opening.
2.60.2 Ventilation System. EH In addition to the standard ventilation system, the EH-60A has a ventilation system which operates in conjunction with the air conditioning
system. The system is controlled from the ECS control
panel on the upper console (Figure 2-7). When the AIR
COND switch is placed in the FAN position, fresh air

TM 1-1520-237-10

is drawn from outside the helicopter into the plenum chamber, mixed with inside air and circulated through the helicopter.
2.60.3 Normal Operation.
1. APU, rotor or external power - Operating.
2. VENT BLOWER switch - ON.
2.61 Air Conditioner System. EH The vapor-cycle
system (air conditioner) cools the cabin and cockpit areas.
It consists of a heli-rotor compressor, evaporator, condenser, associated valves, protective pressure and temperature switches, a filter, service valves, a liquid indicator and
an electrical control system. A sight glass in the liquid line
gives an indication of refrigerant liquid servicing level,
when the system is operating. The temperature controller
assembly, in the aft cabin, processes the input signals from
the temperature selection rheostat in the cockpit and the
cabin temperature sensor, and provides the power to the hot
gas bypass valve solenoid. The electrical control box, in the
transition section, contains the relays, time delays, elapsed
time meter and fault indicators for the vapor-cycle system.
The control box routes the power to the electrical components. Inputs from the remote control and temperature controller are channeled to their respective electrical interface
in the control box. Across the front of the enclosure are
four fault indicators HI and LO PRESS, and HI and LO
TEMP, which are tripped to indicate red when a fault is
received. These indicators provide visual signals of a fault
occurring, even if it is only temporary, and they can be

manually reset for reuse by pressing in the fault indicator.
The air conditioner system is protected to prevent evaporator freezing. The system may be operated at any ambient
temperature without causing damage, shown in Table 2-2.
Power to operate the air conditioner system is provided
from the No. 2 ac primary bus and controlled from the No.
1 dc primary bus through a breaker marked ECS CONTR
. Control of the air conditioner is through the ECS control
panel on the upper console (Figure 2-7). The panel contains
a temperature control rheostat with an increasing arrow indicator to COOL, two mode selection switches marked
COOL-OFF-FAN and HTR-OFF-ON. The temperature
control rheostat is used with the COOL switch to set the
desired cabin temperature. Placing switch to COOL will
cause AIR COND ON advisory light to illuminate. Selection of the COOL mode on the cockpit AIR COND control panel starts a phased sequence of events leading to full
operation of the air conditioner system. To prevent a sudden surge in 115 vac power, the major electrical components are started at spaced intervals.
2.62 AUXILIARY HEATER SYSTEM.

Incorporated in the air conditioner plenum chamber is an
auxiliary heating system to supplement the bleed air heater.
The electrically operated heater is controlled by a switch on
the upper console ECS control panel marked HTR ON &
OFF. The heater element will operate continuously as long
as the switch is ON. With the HTR switch ON and the
AIR COND switch placed in the FAN position the CABIN
HEAT ON advisory light will illuminate. An overtemperature protection is provided at 205°F if there is a heater
malfunction.
Table 2-2. Air Conditioning System Power Source Priority EH

POWER SOURCE

AIR CONDITIONING SYSTEM OPERATION

APU Generator (Aircraft on
Ground)

vAir Conditioning Interrupted if:
(1) Backup Pump is On.
or
(2) Windshield Anti-Ice is On.
vWindshield Anti-Ice Interrupted when Backup Pump is On.

APU Generator (Aircraft in
Flight)

vAir Conditioning Interrupted while Aircraft in Air.
vWindshield Anti-Ice Interrupted when Backup Pump is On.

Dual Main Generator (No. 1 and No.
2) (Aircraft in Flight or on Ground)

vAir Conditioning, Backup Pump, and Windshield Anti-Ice can
Operate Simultaneously.

Single Generator or External AC Power
(Weight on or off Wheels)

vAir Conditioning Interrupted If:
(1) Backup Pump is On.

Change 3

2-63

TM 1-1520-237-10

Section XI ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEMS
2.63 ELECTRICAL POWER SYSTEMS.
Alternating current (ac) is the primary source of power.
The primary electrical system consists of two independent
systems, each capable of supplying the total helicopter
power requirements. The prime source of each system is a
115/200 vac generator. A subsystem feeds two independent
ac primary buses and an ac essential bus. A portion of each
ac primary bus load is converted to 28 volts direct current
(vdc) by two 200 ampere ac/dc converters. The 28 vdc is
distributed by two independent dc primary buses and a dc
essential bus. Emergency power is provided by a generator
driven by the auxiliary power unit (APU). The APU generator is capable of supplying all flight-essential ac and dc
bus loads. In addition, the APU generator can supply power
to the blade deice system (when installed) if one main generator should fail. Should a second generator fail, the blade
deice load will be dropped and the APU generator will
power the remaining ac bus loads. An electric power priority feature allows either the No. 1 or No. 2 main generator
to automatically supersede the APU generator, which, in
turn, automatically supersedes external power. A 24-volt
battery provides backup dc power.
2.64 DC POWER SUPPLY SYSTEM.
Primary dc power is obtained from two converters
(transformer-rectifiers) with a battery as the secondary
power source. There is no external dc power connector
(Figure 2-19).
2.64.1 Converters. Two 200-ampere converters, each
normally powered by the No. 1 and No. 2 ac primary buses
respectively, turn ac power into dc power and reduce it to
28 volts. The converter output is applied to the No. 1 and
No. 2 dc primary buses whenever ac power is applied to the
ac primary buses. If one converter’s output is lost, the converter load will be transferred to the operating system, and
a caution light, marked #1 CONV or #2 CONV will go on.
Power to light the caution light is provided by the battery
bus through a circuit breaker marked, AC CONV WARN.
2.64.2 Battery.
a. A 24-volt dc 5.5 ampere hour 20-cell nickel cadmium
(nicad) battery provides secondary or emergency dc power.
The battery is in the cabin section behind the copilot. It
supplies dc power to the battery and battery utility buses
(Figure 2-19) for operating dc essential equipment during
primary dc malfunction. Power to the battery bus is controlled by the BATT switch on the upper console. It has
marked positions OFF and ON. The battery utility bus is
2-64

Change 4

connected directly to the battery. During No. 1 and No. 2
dc primary source malfunction, the dc essential bus is powered by the battery bus as long as the battery is at least 35%
charged and the BATT switch is ON. When only battery
power is available, the battery life is about 22 minutes day
and 14 minutes night for a battery 80% charged. The BATT
switch should be ON when either external power, APU
generator or main generator power is applied to the helicopter. This will recharge the battery. When the battery is
the sole source of dc power, the BATT switch should be
turned OFF immediately upon obtaining a BATT LOW
CHARGE caution light. A malfunction of both dc primary
sources will light caution lights marked #1 and #2 CONV.
If the BATT switch is left ON, the battery will be completely discharged in less than 3.5 hours. If the maintenance
light and both cockpit utility lights are left on, the battery
will be completely discharged in less than 7 hours. Power
to light the caution light is from the battery bus through a
circuit breaker marked BATT & ESNTL DC WARN EXT
PWR CONTR.
b. A 24-volt dc 9.5 ampere hour sealed lead acid battery
(SLAB) provides secondary or emergency dc power. The
battery is in the cabin section behind the copilot. It supplies
dc power to the battery bus, battery utility bus and dc essential bus (Figure 2-19) for operating dc essential equipment during primary dc malfunction. Power to the battery
bus is controlled by the BATT switch on the upper console.
It has marked positions OFF and ON. The battery utility
bus is connected directly to the battery. The dc essential
bus is powered by the battery bus as long as the BATT
switch is ON. When only battery power is available, the
battery life is about 38 minutes day and 24 minutes night
for a battery 80% charged. The BATT switch should be
ON when either external power, APU generator or main
generator power is applied to the helicopter. This will recharge the battery. When the battery is the sole source of dc
power, the BATT switch should be turned OFF immediately upon obtaining a BATT LOW CHARGE caution
light. This is done so that battery power can be conserved
for an APU start. A malfunction of both dc primary sources
will light caution lights marked #1 and #2 CONV. If the
BATT switch is left ON, the battery will be completely
discharged in less than 6 hours. Power to light the BATT
LOW CHARGE caution light is from the battery bus
through a circuit breaker marked BATT & ESNTL DC
WARN EXT PWR CONTR.
2.64.3 DC Monitor Bus. EH The dc monitor bus is normally energized by the No. 1 and No. 2 converters when
the generators are operating, and is powered by the No. 2
converter when operating from external power (Figure

TM 1-1520-237-10

UH60 ELECTRICAL SYSTEM
MAIN GENERATORS
#1
GENERATOR

115 / 200 VAC
30 / 45 KVA
400 HZ
3 PHASE

#2
GENERATOR

#1 GEN

GCU

GENERATOR CONTROL UNIT
UNDERFREQUENCY PROTECTION (375 5 HZ / 1−3 SEC)
(GROUND OPERATION ONLY ON MAIN GEN GCU’S)
UNDERVOLTAGE PROTECTION (100 5 VAC / 6 1 SEC)
OVERVOLTAGE PROTECTION (125 VAC / TIME VARIES
INVERSELY TO OVERVOLTAGE)
FEEDER FAULT (SHORT CIRCUIT TO GROUND)

GCU
#2 GEN

APU
GENERATOR

GCU
60 AMP
CURRENT LIMITERS
(6 TOTAL)

#1 AC
PRIMARY BUS

APU GENERATOR

#2 AC
PRIMARY BUS

AC ESS
BUS OFF

AC
ESSENTIAL BUS
AC TO DC
CONVERTER

115 / 200 VAC
20 / 30 KVA, 400 HZ
3 PHASE, AIR COOLED

EXT PWR
MON PANEL

EXTERNAL AC
POWER

AC TO DC
CONVERTER

EXTERNAL POWER MONITOR PANEL

CONVERTS 115 / 200 VAC
TO 28 VDC 200 AMPS

#1 CONV

UNDERVOLTAGE PROTECTION
(100−105 VAC / .85−2.55 SEC)
OVERVOLTAGE PROTECTION
(125−130 VAC / 1 .25 SEC)
UNDERFREQUENCY PROTECTION
(370−375 HZ / 1 .25 SEC)
OVERFREQUENCY PROTECTION
(425−430 HZ / 1 .25 SEC)
CORRECT PHASE ROTATION

#2 CONV

#1 DC
PRIMARY BUS

#2 DC
PRIMARY BUS
100 AMP
CURRENT LIMITER

AA0327_1
SA

Figure 2-19. Electrical System (Sheet 1 of 2)

Change 2

2-64.1

TM 1-1520-237-10

DC
ESSENTIAL BUS

BATTERY
FAULT

BATT
LOW CHARGE

INTERNAL TEMP ABOVE
70oC OR CELL
DISSIMILARITY EXISTS

BATTERY FALLS
BELOW
40 %
CHARGE

CHARGER
ANALYZER
20 CELLS
5.5 AMPERE HOUR

BATTERY
UTILITY BUS

BATTERY
BUS

24 VDC
BATTERY

BATTERY SWITCH

DC ESS
BUS OFF
WHEN BATTERY IS
ONLY SOURCE OF
POWER,
DC ESSENTIAL BUS
IS DROPPED IF
BATTERY FALLS
BELOW 35% CHARGE

HELICOPTERS WITH NICAD BATTERY INSTALLED

BATTERY LOW
SENSING RELAY

BATT
LOW CHARGE
BATTERY FALLS
BELOW
23 VOLTS

DC
ESSENTIAL BUS

9.5 AMPERE HOUR

1
BATTERY
UTILITY BUS

BATTERY
BUS

24 VDC
BATTERY

BATTERY SWITCH

DC ESS
BUS OFF

HELICOPTERS WITH SLAB INSTALLED

Figure 2-19. Electrical System (Sheet 2 of 2)

2-64.2

Change 10

AA0327_2B
SA

TM 1-1520-237-10

2-19). If either converter should fail, the bus will be automatically dropped from the system.
2.64.4 Quick Fix Power. EH Mission equipment dc
power is provided from the No. 1 dc primary bus, and is
controlled by Q/F PWR switch on the upper console.
2.64.5 Battery Charger/Analyzer. A charger/analyzer
system restores the battery charge and determines the condition of the battery. The system charges the battery
through a converter whenever ac power is available on the
helicopter and the BATT switch is ON. The analyzer system monitors battery charge and lights a caution light indi-

cating BATT LOW CHARGE when the charge lowers to
35% to 45% of battery capacity. If battery charge continues
to lower, at 30% to 40% of battery capacity, the dc essential bus will be disconnected from the battery. At 35% capacity the battery can provide two APU starts. Another
analyzer circuit monitors battery temperature. When the internal temperature reaches 70°C (158°F), or if a battery cell
dissimilarity condition exists, a caution panel light will go
on, indicating BATTERY FAULT (only on helicopters
equipped with nickel-cadmium batteries). Then the charger/
analyzer should automatically disconnect the battery from
the charging circuit. As a backup, placing the BATT switch

Change 2

2-65

TM 1-1520-237-10

POWER SOURCE FOR JETTISON SYSTEM
WHEN M−139 KIT IS INSTALLED

NO. 1
AC P
RI BU
S
60 HZ AC
CONVERTER
15

WSHLD
WIPER

NO. 1
CONVERTER

IFM

LIGHTS
CPTL GLARE UPPER CABIN
5

SHLD
FLT
AC ESNTL NO. 1
AC
BUS

LWR

NO. 1
ENG

HUD

LWR

CSL

DOME
LEFT
PITOT

CSL

OVSP

REF

SYS

CSL

7.5

SPLY

INST

HEAT

AIR
UTIL SOURCE FUEL BACKUP ESSS JTSN
RECP HEAT/ LOW PUMP INBD OUTBD

CPLT WSHLD
ANTI−ICE

7.5

NO. 1
DC P
RI BU
S

7.5

CABIN START WARN PWR
LIGHTS
NO. 1 ENG
ADVSY CAUT
RETR LDG
WARN
ANTI−ICE
5

7.5

ADVSY CONT
DPLR

IFF

CPTL NO. 1
WSHLD DC

LTS
WARN ANTI−ICE INST
CMPTR CHAFF
CPLT
ADF CMD CSL TRIM
DISP TURN ALTM MODE
2

PWR

7.5

RATE GYRO

SET

BUS
TIE
5

NO. 1 T RTR
GEN SERVO
5

NO. 1
SERVO
5

CNTOR WARN WARN CONTR WARN
DC ESNTL
NO. 2
VHF FM COMM
RDR
BUS

SELECT

FM SCTY SET ALTM WARN

SPLY

POWER SOURCE FOR M−139 SYSTEM
WHEN KIT IS INSTALLED

(SEE NOTE 4)
UH

NO. 1 CIRCUIT BREAKER PANEL
NO. 1
AC P
RI BU
S
NO. 1 FUEL

IINS

CPLT WSHLD
ANTI−ICE

BOOST PUMP
WSHLD
WIPER

NO. 1
CONVERTER

EXT
FUEL
5
LH

CPLT

SEC

FLT
AC ESNTL NO. 1
AC
BUS

LIGHTS
UPPER CABIN
5

CSL DOME
Q/F
LEFT
XFMR PITOT

7.5

SPLY

INST

PWR

HEAT

LWR
5
CSL
ICE
DET
2

NO. 1
ENG
5

ALQ−156
2

φC
OVSP
INU 26 VAC
BATT EQUIP
5

PWR

AIR
FUEL SOURCE
LOW HEAT/
5

IFM
10

BACKUP T / R
PUMP DE−ICE
.5

ECS

NO. 1
DC P
RI BU
S

PWR
WARN START
PWR
PWR
LIGHTS
NO. 1 ENG
ADVSY CAUT
RETR LDG
WARN
ANTI−ICE
5

7.5

φB

φA

28V

IINS

DE−ICE

DET

IFF

ADF

7.5

ADVSY CONT

CONTRLR

CPTL NO. 1
WSHLD DC

PWR
LTS
WARN ANTI−ICE INST
CPLT
Q / F CMPTR CHAFF
DISP TURN ALTM MODE
EQUIP TRIM
25
PWR

7.5

DISP RATE GYRO

2
SEL

BUS
TIE
5

NO. 1 T RTR
GEN SERVO
5

NO. 1
SERVO
5

ESSS
JTSN
7.5

CNTOR WARN WARN CONTR WARN INBD
DC ESNTL ESSS
NO. 2
JTSN
VHF FM COMM
RDR
BUS
5

FM SCTY SET ALTM WARN

7.5

SPLY OUTBD

NO. 1 CIRCUIT BREAKER PANEL

NOTES
1. ON HELICOPTERS PRIOR TO SERIAL
NO. 81−23579 NOT MODIFIED BY MWO
55−1520−237−50−25.
2. ON HELICOPTERS PRIOR TO SERIAL
NO. 81−23579 MODIFIED BY MWO
55−1520−237−50−25 AND HELICOPTERS
SERIAL NOS 81−23579 AND
SUBSEQUENT.
3. CIRCUIT BREAKERS WITH ASTERISK ( )
ARE ON HELICOPTERS SERIAL NOS
82−23748 AND SUBSEQUENT.
4. ON HELICOPTERS MODIFIED BY MWO
1−1520−237−50−62, HUD.
5. ON HELICOPTERS SERIAL NOS
97−26744 AND SUBSEQUENT.

AA0353_1E
SA

Figure 2-20. DC and AC Circuit Breaker Panels (Typical) (Sheet 1 of 4)

2-66

Change 1

TM 1-1520-237-10

S
RI BU
AC P
NO. 2
HEAT & VENT

S
RI BU
DC P
NO. 2

WINDSHEILD
ANTI−ICE
5

SERVO
5

NO. 2
DC
5

GEN

BUS
TIE

BATT

7.5

SPLY SELECT

VENT

CONTR TRIM

7.5

PWR

TRIM

UTIL RECP

PILOT WSHLD
ANTI−ICE

7.5

FIRE

CMPTR

7.5

WARN CONTR INST WARN CNTOR CHGR WARN
LTS CONTR PWR CONTR SPLY
DC ESNTL
PILOT
MAIN
BUS MODE ALTM HEAT
VHF
IRCM CMPTR STAB SPEED XMSN POS
STAB
50

7.5

NO. 2 CONVERTER
CTR WSHLD
ANTI−ICE

PILOT CTR EXTGH
NO. 2 ENG
AC ESNTL
ANTI−ICE
WARN START CARGOHOOK BUS
BATT
5

STAB

5
CHGR

NO. 2
ENG

HEAT CONTR INST

LIGHTS
ANTI
PLT

FORM

NON CARGO

OVSP

COLL

FLT

HOOK

AC ESNTL BUS
HSI

CIS

SAS

LTS

RT
PITOT

26 VAC
2

COMP

AMPL STAB IND INST

CONTR PLT / CPLT

DPLR

VSI

AUTO AC ESNTL

PLT

CPLT

XFMR BUS WARN

NO. 2 CIRCUIT BREAKER PANEL

NO. 2 EXTD
RANGE PUMP

AUX HTR

AUX FUEL QTY

15
BLOWER

15
NO. 2 FUEL

A
C

ICE−DET

A
C

AUX HTR

A
C

ICE−DET

BOOST PUMP

NO. 2 PRI BUS

BLOWER
NO. 2 FUEL

A
C

NO. 2 PRI BUS
BOOST PUMP

RESQ HST

NO. 2 LTR

CONTROL

LTS

DE−ICE ICE−DET

D
NO. 2 XFER C

7.5

CNTRLR

RESQ HST

CONTROL

NO. 2 LTR

CONTROL

LTS

7.5

CONTROL

CNTRLR

NO. 1 XFER
5

AUX HTR

NO. 1 LTR
CONTROL

D
C

CONTROL

D
EL NO. 2 XFER C
5

EXT FUEL NO. 1 LTR NO. 1 XFER

D
C

CONTROL

EXT FUEL NO. 2 XFER C

DE−ICE ICE−DET

LTS

CONTROL

D
C

CONTROL

NO. 1 FUEL

LTS

(ON HELICOPTERS
EQUIPPED WITH
AUXILIARY CABIN
HEATER)

BOOST PUMP
DE−ICE PWR

DE−ICE PWR
20

A
C

TAIL ROTOR

NO. 1 EXTD
RANGE PUMP

(SEE NOTE 1)

D
C

NO. 1 PRI BUS

AUX HTR
5

NO. 1 PRI BUS

H CONTROL

= ES

A
C

(SEE NOTES 2 AND 3)

MISSION READINESS
CIRCUIT BREAKER PANELS (CABIN)

AA0353_2A
SA

Figure 2-20. DC and AC Circuit Breaker Panels (Typical) (Sheet 2 of 4)

2-67

TM 1-1520-237-10

S
RI BU
AC P
NO. 2
HEAT & VENT

ON
DC M
S
RI BU
DC P
NO. 2

WINDSHIELD
ANTI−ICE
5

TACAN
2

SERVO
5

NO. 2
DC
5

GEN

BUS
TIE

BATT

7.5

FIRE

UTIL
RECEPT ALQ−162
7.5

CTR EXTGH CABIN
PILOT
NO. 2 ENG
ANTI−ICE
WARN START ECS
5

7.5

WARN CONTR INST WARN CNTOR CHGR WARN
LTS CONTR PWR
SEC MON DC ESNTL
DC MON
PILOT
MAIN
BUS
BUS MODE ALTM HEAT
VHF
IRCM CMPTR STAB SPEED XMSN POS
BUS
5
CONTR

SPLY SELECT

VENT

CONTR TRIM

7.5

PWR

TRIM

BUS

7.5

CMPTR

N0. 2 CONVERTER
CTR WSHLD
ANTI−ICE

UTIL RECP

PILOT WSHLD
ANTI−ICE

7.5

28V
EXT
FUEL

AUX
FUEL

NO. 2 FUEL

TACAN

RH
IR

LTS

SPLY

LTS

CONTR
BOOST PUMP
AC ESNTL
RIGHT
BUS
BATT PITOT
7.5

SPLY CHGR

STAB
5

LIGHTS
ANTI
PLT

FORM
5

LV
HV
COLL
FLT
NO. 2
AUX FUEL
AC
QTY
ENG
5

HEAT CONTR INST

5
OVSP

φA

NON
5
FLT
ALQ−162
3

φB

φC

EH
PILOT’S CIRCUIT BREAKER PANEL

AA0353_3B
SA

Figure 2-20. DC and AC Circuit Breaker Panels (Typical) (Sheet 3 of 4)
OFF removes input power to the charger/analyzer. By
placing BATT switch OFF, the increasing temperature may
be checked.
2.64.6 Battery Low Sensing Relay. On helicopters
equipped with the sealed lead acid battery the system
charges the battery through the battery charging relay with
one or both converters on. A caution light indicating BATT
LOW CHARGE lights when voltage on the battery utility
bus drops below 23 vdc.
2.64.7 DC and AC Circuit Breaker Panels. The circuit breaker panels (Figure 2-20) protect the power systems. One is above and to the rear of each pilot and copilot,
one is on the lower console, and two are on the upper
console . The ac essential bus contains one additional panel.
The circuit breakers provide both ac and dc protection.
Popping of a circuit breaker indicates too much current is
being drawn by a component in the circuit that is powered
through the circuit breaker. Unnecessary recycling of circuit breakers, or using circuit breakers as a switch should
not be done.

2-68

Change 9

2.65 AC POWER SUPPLY SYSTEM.
A primary ac power system (Figure 2-19) delivers regulated three phase, 115/200 vac, 400 Hz. Each system contains a 30/45 kilovolt-ampere generator mounted on and
driven by the transmission accessory gear box module, a
current transformer, a generator control unit, and current
limiter, all of which are interchangeable. System outputs
are applied to the No. 1 and No. 2 ac primary buses. Caution lights will go on, indicating #1 GEN or #2 GEN whenever generator output is interrupted. Another caution light
goes on, indicating AC ESS BUS OFF when there is no
power to the ac essential bus. Individual generator controls
are provided on the upper console (Figure 2-7), with
marked positions of TEST, OFF/RESET, and ON. A generator main bearing caution system is installed on each
main generator to light a caution light, marked #1 GEN
BRG or #2 GEN BRG, to indicate a worn or failed bearing. The caution light will remain on until power is removed. The auxiliary bearing will allow 10 additional hours
of operation after the light goes on. Therefore, it should not
be a cause for mission abort. Power to operate the caution

TM 1-1520-237-10

POWER SOURCE FOR EMERGENCY
JETTISON SYSTEM WHEN M−139
KIT IS INSTALLED

DC ESNTL BUS

DC ESNTL BUS
ICS

NO. 1 VOR / ILS CHIP
2

PILOT COPILOT VHF FM

COMM SCTY SET UHF
NO. 1 FM UHF AM AM
2

DET

FUEL
PUMP

ESSS
JTSN

7.5

STAB

CONTR OUTBD

CARGO PILOT
HOOK TURN

PNL

7.5

EMER

DETR

ENG

SENSE

SPLY

SAS
5

7.5

CONTR SHEAR

BATT
BUS

PWR

CAUT / BACKUP HOIST ESSS
HYD CABLE JTSN
ADVSY

7.5

ESNTL
DC

FIRE DET
NO.1
NO.2

NO. 1
TAIL
ENG WHEEL
5

CONTR SRCH

PNL

PWR

CONTR

BOOST START LOCK

INBD

LIGHTS
SEC

UPPER CONSOLE CIRCUIT BREAKER PANEL

ESNTL BUS
DC

B 50
A
T SPLY
T
B
U
S

AC &
5

BATT &
ESNTL DC FUEL B BATT
WARN PRIME A BUS
5

T
T

ESNTL BUS
FIRE

B 50
A
T SPLY
T

CONV EXT PWR BOOST U CONTR EXTGH
T
WARN CONTR
I UTIL
APU
APU
LTS

B
U
S

CONTR
INST

FIRE
DET

B
U
GEN
GPS S CKPT
CONTR ALERT

AC &

CONTR
INST

BATT &
ESNTL DC FUEL B BATT
WARN PRIME A BUS
5

T
T

FIRE
5

CONV EXT PWR BOOST U CONTR EXTGH
T
WARN CONTR
UTIL
I
APU
APU
LTS

CONTR
INST

FIRE
DET

GEN
CONTR

GPS

B
U
CKPT S

5
CONTR
INST

(SEE NOTE 5)

BATTERY AND BATTERY UTILITY
BUS CIRCUIT BREAKER PANEL

AC ESNTL BUS
STAB

HSI

CIS

SAS

CONTR

PLT / CPLT

AMPL

26 VAC
2

STAB

IND INSTL

COMP

VSI

AUTO

PLT

CPLT

XFMR

AC ESNTL

BUS WARN

AC ESSENTIAL BUS
EH
AA0353_4D
SA

Figure 2-20. DC and AC Circuit Breaker Panels (Typical) (Sheet 4 of 4)

Change 1

2-69

TM 1-1520-237-10

system is provided from the No. 1 and No. 2 dc primary
buses, through circuit breakers, marked NO. 1 GEN
WARN and NO. 2 GEN WARN, respectively.

be off at any time either No. 1 generator or No. 2 generator
is supplying power. The generator control switch on the
upper console (Figure 2-7), has marked positions of TEST,
OFF/RESET, and ON.

NOTE
NOTE
When the GEN BRG caution light remains
on for more than 1 minute, make an entry on
the DA Form 2408-13-1.
2.65.1 Generator Control Units (GCU). The GCUs
monitor voltage from the No. 1, No. 2 and APU generators
and take the generator(s) off-line where malfunctions occur.
Underfrequency protection is disabled in flight by the
WOW switch.
2.65.2 AC Secondary Bus. EH The ac secondary bus
is powered by the No. 1 and No. 2 generators when they are
operating and their outputs are acceptable (Figure 2-19).
Current limiters protect the system from excessive current
draw. If the No. 1 and No. 2 generators are off, the APU
generator will supply the ac secondary bus if the output is
acceptable, the backup hydraulic pump is off, the blade
deice is off, and the weight of the helicopter is on the
wheels. The ac secondary bus can also receive power from
external power when the weight of the helicopter is on the
wheels, and the No. 1, No. 2 and APU generators are off,
and the backup hydraulic pump is not operating.
2.66 AUXILIARY AC POWER SYSTEM.
An auxiliary ac power system (Figure 2-19), is a backup
ac power source, providing electrical power for ground
checkouts. The system consists of a 115 vac three-phase,
400 Hz 20/30 kVA, air-cooled generator mounted on and
driven by the APU, a current transformer and a generator
control unit. If the primary ac generators are not operating,
the auxiliary ac power output will be applied through contactors to the No. 2 ac primary bus and through contactors
and current limiters to the No. 1 ac primary bus. An advisory light on the caution/advisory panel will go on, indicating APU GEN ON when the APU generator is operating,
and the APU generator switch is ON. APU GEN ON light
will be on only when supplying power to the system, it will

2-70

Change 3

If the APU generator is the sole source of ac
generated power, all equipment may be operated, except that when the backup pump is
on, the windshield anti-ice and EH air conditioner are prevented from being used.
2.66.1 Generator Control Switches. Generators are
controlled by a three-position generator switch on the upper
console (Figure 2-7). The switch ON position energizes the
generator and permits connection of generator ac output to
the ac loads. TEST permits you to test the generator ac
output without connecting to the generator loads. OFF/
RESET deenergizes the generator and permits generator
recycling if the generator is disabled and disconnected from
its loads. The control switch is manually placed to RESET
and then back to ON.
2.66.2 External AC Power System.

CAUTION

Do not connect a source of dc power to
the external ac connector.
An external ac power connector, on the right side of the
helicopter (Figure 2-1), accepts ground source of 115 vac,
three-phase, 400 Hz power. The system is controlled by a
switch on the upper console (Figure 2-7), marked EXT
PWR-RESET-OFF and ON. External power will be introduced into the system if acceptable external power is connected, the EXT PWR switch is ON, and no other generating source is operating. An advisory light on the caution/
advisory panel will go on, indicating EXT PWR
CONNECTED, whenever external power is connected to
the helicopter.

TM 1-1520-237-10

Section XII AUXILIARY POWER UNIT
2.67 AUXILIARY POWER UNIT (APU) SYSTEM.
The auxiliary power unit system (Figure 2-21) consists
of an auxiliary power unit (APU), accessories, controls, a
monitoring system, and a starting system. The APU system
provides pneumatic power for main engine starting and
cabin heating, and electrical power for ground and emergency in-flight electrical operations.
NOTE
The APU is not qualified for normal inflight
use.
APU system accessories include a prime/boost pump,
hydraulic accumulator, hydraulic handpump, hydraulic start
motor, and ac generator. The hydraulic accumulators and
handpump, in the aft midsection cabin ceiling (Figure 2-5),
provide the hydraulic pressure for driving the APU starter.
If the APU does not start, the hydraulic accumulator can be
recharged by pumping the hydraulic handpump. The hydraulic utility module and backup pump, on the left forward
deck within the main rotor pylon, will automatically recharge the depleted hydraulic accumulator for the next APU
start. The APU controls are in the cockpit on the upper
console. Indicator lights on the caution/advisory panel provide cockpit monitoring of the APU. An indicator panel in
the cabin will indicate reason for APU shutdown on BITE
indicators. The BITE indicators are incorporated in the
APU electronic sequence unit (ESU), and will indicate reasons for APU shutdown. Those indicators can be monitored
during APU operation without interrupting normal operating systems. During a start, the ESU compares input signals
from speed, time, and temperature sensors on the APU to
specified values stowed in the ESU memory, and performs
functional steps as a result of the comparison. The system
also provides for APU protective shutdown in case of turbine overspeed, underspeed, high exhaust temperature, low
oil pressure, or loss of electrical power or sequence failure.
Each major sequence step will have a visual indication of
go/no-go. The ESU samples predetermined parameters of
exhaust temperature, turbine speed and oil pressure. If any
one of the predetermined values are exceeded, the APU
will shut down, and appropriate BITE indication is made.
On helicopters modified with improved ESU, if a momentary malfunction occurs (i.e., a power interruption other
than switching of the APU CONTR switch) the APU will
shut down and the APU CONTR switch must be placed at
OFF and then back ON, to restart the APU. There is also
an output signal to the caution/advisory panel to turn on the
APU ON advisory light, indicating the APU is operating.

Power to operate the APU and ESU is provided from the
battery bus through a circuit breaker marked APU CONTR
INST.
2.68 APU.
The auxiliary power unit (Figure 2-21) consists of a gas
turbine shaft power section, a reduction gear drive, and
appropriate controls and accessories. The accessory gear
box provides an axial pad with a 12,000 rpm output drive
for the APU ac generator, rpm pad for mounting the APU
start motor, rpm drive pad for the APU fuel assembly. A
magnetic pickup mounted on the accessory gear box senses
engine speed. The APU is lubricated by a self-contained oil
system. Refer to Figure 2-25 for servicing.
2.68.1 APU Controls. The APU control, on the upper
console (Figure 2-7), consists of a CONTR switch and an
APU fire extinguisher T-handle. The APU CONTR switch,
with marked positions OFF and ON, controls the operation
of the APU. Placing the switch ON starts the APU and
allows it to operate. The APU is off when the switch is
OFF. The APU FAIL caution light will be on any time the
APU automatically shuts down. The APU OIL TEMP HI
caution light is on when APU oil temperature is above normal range. During ground operation at high ambient temperatures the APU OIL TEMP HI caution light may go
on. If this occurs, the APU should be shut down immediately to prevent damage. After a 30-minute cooling period,
the oil level should be checked. If OK, the APU may be
restarted. The control system receives electrical power from
the battery bus through a circuit breaker marked APU
CONTR INST on the lower console. When illuminated,
the APU T-handle warns the pilot/ copilot of a fire in the
APU compartment. When the T-handle is pulled, it turns
off fuel to the APU, sends a stop signal to the ESU, arms
the fire extinguisher system, and sets the extinguisher direction control valve to the APU. During APU starts using
battery power, if the fire extinguisher is required, FIRE
EXTGH RESERVE must be used. The T-handle microswitch receives electrical power from the battery utility
bus through a circuit breaker marked FIRE EXTGH on the
lower console circuit breaker panel.
2.68.2 APU Fuel Control System (Helicopters
equipped with (T-62T-40-1 APU). This system consists
of a fuel pump and a control assembly. The fuel pump is
protected by a filter. Fuel pump output flow passes through
another filter before entering the control assembly. A governor and flow metering valve controls fuel flow to the
engine during ignition, permitting automatic starting under

2-71

TM 1-1520-237-10

FUEL CONTROL
ENCLOSURE

FIREWALL
FIRE DETECTOR

ELECTRICAL
CONNECTOR

HYDRAULIC
START MOTOR

FIREWALL

GENERATOR
OIL FILLER PORT
AND DIPSTICK

OIL LEVEL
SIGHT GAGE

BLEED−AIR
PORT

AA0519
SA

Figure 2-21. Auxiliary Power Unit (APU) (Typical)
all ambient conditions, and controls the turbine at a constant speed once it has accelerated to operating speed. An
electronic speed sensing device provides automatic fuel
flow, ignition, and operation of the APU.
2.68.3 APU Fuel Control System (Helicopters
equipped with GTC-P36-150 APU). The fuel control
system includes a fuel pump and metering section. The fuel
pump is protected by an integral inlet filter. Fuel pump
output flow passes through a filter screen before entering
the metering assembly. Fuel pump discharge pressure is
limited by an ultimate relief valve which, when activated,
bypasses fuel flow back to the pump inlet. Fuel metering is
accomplished by the torque motor metering valve as a
function of an electrical signal from the electronic sequence
unit (ESU). For accurate fuel metering, a constant, pressure
drop across the metering valve is maintained by the differential pressure regulating valve. The fuel solenoid valve is
energized by the ESU following the initiation of APU start.
This allows fuel to flow to the engine. The fuel control
assembly subsequently provides fuel according to a preprogrammed schedule to effect efficient acceleration. The
fuel solenoid valve will close completely without visible
leakage from the minimum operating fuel pressure to 110%
of the maximum operating fuel pressure.

2-72

2.68.4 APU Fuel Supply System. APU fuel is supplied to the APU from the left main fuel tank. The FUEL
PUMP switch must be at APU BOOST for all APU operation, except engine priming. The APU prime/boost shutoff valve is a two-position, open-closed unit mounted on
the APU compartment firewall where it also functions as a
firewall shutoff valve. The valve is pilot-operated from the
upper console FUEL PUMP switch as well as by the FIRE
EXTGH APU T-handle. If the APU does not start and the
APU ACCUM LOW advisory light is not on, the manual
override lever on the accumulator manifold should be
pulled to attempt another start, and held until the APU has
reached self-sustaining speed.
2.69 ACCUMULATOR RECHARGE.
The accumulator recharge cycle starts when the APU
has reached operational speed and the APU-driven generator comes on the line. The pressure switch for the accumulator causes the APU ACCUM LOW light to go on and
the backup system pump to develop pressure. The APU
accumulator pressure should be at least 2800 psi before
attempting an APU start. The accumulator is recharged
from the backup pump which runs for 90 seconds after the
accumulator low-pressure switch is actuated. When the

TM 1-1520-237-10

winterization kit is installed, an additional identical accumulator is installed in parallel with the original accumulator. Discharge and recharge of the added accumulator is the
same, except a 180-second recharge cycle for the two accumulators will take place when the accumulator pressure
switch senses low accumulator pressure. Both accumulators
are charged or discharged simultaneously. If the accumulators do not fully charge during the first 180 seconds of the
backup pump operating cycle, the pump will continue to
operate in 180-second segments, or until the BACKUP
PUMP PWR circuit breaker is pulled, or 115 vac power is
removed. The backup system pump shuts down after recharge, unless required for other purposes. Should the accumulator pressure drop, the backup system pump restarts
to replenish the accumulator charge. The rate of accumulator charge is limited to protect the backup system from

possible depletion due to ballistic damage to the APU start
system. Should the APU not start, the accumulator may be
recharged by these methods, after the APU CONTR switch
is OFF. An electric ground cart powering the backup hydraulic pump or a hydraulic ground cart connected to the
backup hydraulic system through the ground test quickdisconnects or by using the handpump in the aft upper
cabin. The APU CONTR switch should not be turned ON
again or the BATT switch turned OFF until after the ESU
BITE indicators have been checked. The handpump may
also be used to top off the accumulator charge if the charge
has dropped due to a low temperature condition. A pressure
gage mounted in the aft cabin (Figure 2-5) indicates the
charge. Check valves prevent draining of the accumulator
charge through the system.

2-73

TM 1-1520-237-10

Section XIII LIGHTING
2.70 INTERIOR LIGHTING.
The interior lighting system consists of cockpit dome
lights, utility lights and cabin dome lights (Figure 2-4).
NVG blue-green lighting can be selected for the cockpit
dome, instrument panel glare shield, utility lights and cabin
dome lights.
2.70.1 NVG Lighting System. The NVG lighting system consists of interior NVG blue-green lighting. Exterior
lighting consists of cargo hook well area electroluminescent lighting, infrared formation and position lights, and
attachable/detachable controllable searchlight filter. A dimming feature is incorporated in the searchlight system to
provide dimming through the collective SRCH LT PUSH
ON - OFF, BRT, DIM switch. The position and formation
lights have IR emitters installed within close proximity to
the regular installed lights to enhance outside viewing with
night vision goggles.
2.70.2 Cockpit Floodlights. Two blue-green and two
white cockpit floodlights are on the overhead cockpit floodlight panel, marked BLUE, OFF and WHITE (Figure 2-7).
Power is supplied from the dc essential bus through a circuit breaker marked LIGHTS SEC PNL. Six lights installed in the instrument panel glare shield provide secondary lighting for the instrument panel. The lights are
mechanically dimmed by a control on the upper console
labeled GLARESHIELD LIGHTS with marked positions
OFF and BRT. Power to operate the glare shield lights is
provided from the No. 1 ac primary bus through a circuit
breaker, marked LIGHTS GLARE SHLD.
2.70.3 Flight Instrument Lights. Instrument lights are
grouped into flight instrument and nonflight instruments.
The flight instrument lights are divided into pilot’s and copilot’s. Lights are controlled by individual rotary intensity
controls (Figure 2-7), marked INSTR LT PILOT FLT,
OFF and BRT, and CPLT FLT INST LTS, OFF and
BRT. The nonflight instrument lights operate in the same
manner as the flight instrument lights. The nonflight lights
intensity is controlled by a rotary control, marked INSTR
LT NON FLT, OFF and BRT. Instrument lighting is provided by instrument bezels with NVG lights. The radar
altimeters lighting incorporates dimming controls on the
instrument panel, marked RAD ALT DIMMING for pilots
radar altimeters (Figure 2-9). The vertical instrument display system has NVG information panel lighting to make
those instruments compatible with the NVG system. Power
to operate the instrument lights is provided by the No. 2 ac
primary bus through circuit breakers marked LIGHTS

2-74

Change 5

PLT FLT and LIGHTS NON FLT, and No. 1 ac primary
bus, through a circuit breaker marked LIGHTS CPLT
FLT.
2.70.4 Lighted Switches Dimmer. A dimmer control
labeled LIGHTED SWITCHES (Figure 2-7) is provided
on the upper console to reduce illumination level of the
following panel lighted switches: Pilot and copilot MODE
SEL, TAILWHEEL LOCK, CIS MODE SEL, AUTO
FLIGHT CONTROL and NO. 1 and NO. 2 FUEL
BOOST PUMP on lights. The caution/advisory panel must
be in DIM mode.
2.70.5 Upper and Lower Console Lights. NVG
lights for the upper console, cockpit flood secondary lights,
engine control quadrant, flight control panel, miscellaneous
switch panel, boost pump control panel, ESSS related panels, range extension fuel management panel, retransmission
control and rescue hoist panels, and compass are illuminated from the No. 1 ac primary bus through dimmer controls marked CONSOLE LT UPPER and LOWER. Circuits are protected by circuit breakers marked LIGHTS
UPPER CSL and LIGHTS LWR CSL. All other lower
console panels are illuminated by the lower console auxiliary utility light next to the copilot’s seat.
2.70.6 Utility Lights. All utility lights are dual (blue/
green-white) (Figure 2-4). Two portable cockpit utility
lights with coiled cords are attached to the upper console by
removable brackets, one on each side of the console. The
lights may be adjusted on their mountings to direct the light
beams or they may be removed and used portably. All utility lights are controlled by a rheostat or a pushbutton on the
end of each casting. The lens casting of the lights may be
turned to change from white to blue/green and/or spot to
flood. An auxiliary utility light, located at the right rear of
the copilot’s seat, is used to illuminate some panels on the
lower console for night flight. On helicopters equipped with
a transition equipment bay, a utility light is installed on the
bay shelf to provide bay lighting EH . The utility lights operate in the same manner as above. Make certain cockpit
utility lights are OFF when not in use. The utility lights
operate from the battery utility bus through a circuit breaker
marked UTIL LTS CKPT. On helicopters 97-26744 and
subsequent, utility lights operate from the battery bus
through a circuit breaker marked UTIL LTS CKPT.
2.70.7 Cabin Dome Lights. Three dome lights are provided for cabin lighting (Figure 2-5). Control of cabin lights
is from the upper console by a control marked CABIN

TM 1-1520-237-10

DOME LT (Figure 2-7) with intensity control and a light
color selector switch. The intensity control has marked positions OFF and BRT, and the light level control may be
adjusted to any position between the two extremes. The
light color selector switch has marked positions WHITE,
OFF, and BLUE. To place the switch from OFF to
WHITE, the switch must first be pulled out to clear a detent. This prevents accidentally placing the switch to
WHITE. Dimming control for the cabin dome lights is
from a control on the left side of the pilots seat (Figure
2-4), marked CABIN DOME LT, with marked positions
OFF and BRT. Power to operate the cabin dome light system is provided from the No. 1 ac primary bus through a
circuit breaker marked LIGHTS CABIN DOME.
2.70.8 Maintenance Light. A portable 20 watt floodlight, in the cabin at the crew chief station is used by the
crew for maintenance work. The light has a 20-foot cord,
allowing its use within the cabin and around the main transmission. A switch on the rear end of the light with marked
positions, DIM, OFF, and BRIGHT, controls the light intensity. Another maintenance light receptacle, in the aft
tailcone, allows the light to be used around the tail section.
Power to operate the light is from the battery utility bus
through a circuit breaker marked UTIL LTS CKPT. The
maintenance light is stowed in a bag at the back of the
pilot’s seat. Power to operate the maintenance lights is provided from the battery utility bus through a circuit breaker,
marked UTIL LTS CKPT. Make sure the maintenance
and cockpit utility lights are OFF when not in use.
2.71 EXTERIOR LIGHTS.
2.71.1 Searchlight.

CAUTION

Landing and searchlight have less than
one foot ground clearance when extended.
Use caution when ground taxiing over
rough terrain when landing light and/or
searchlight are extended.
UH The searchlight (Figure 2-1) is mounted on the right
bottom of the nose section, and is controlled from either
collective pitch stick. The 150 watt light can be moved
forward through a 120° arc from the stow position. It can
also be turned 360° in either a right or left direction on its
axis. The light is operated by a switch labeled SRCH LT
ON, OFF, BRT, and DIM (Figure 2-4). Directional control
of the light is provided through the four-position searchlight control switch, labeled EXT (extend), RETR (retract),
L (left), and R (right). When the SRCH LT switch is

placed ON, the lamp will go on, arming the control switch.
Placing the control switch to EXT causes the light beam to
move forward at a rate of about 12° per second. If the
switch is placed to OFF the light will extinguish. To retract
the searchlight, place the switch to RETR. Refer to Chapter 5 for extend/retract limitations. An infrared filter can be
installed on the controllable searchlight to enhance viewing
objects outside the helicopter when wearing the night vision goggles. With the IR filter installed, maximum wattage
lamp to be used is 250 watt. An OUTPUT switch on the
searchlight dimmer under the pilot’s seat, is placed at
NORM when dimming feature on searchlight is desired.
When in BYPASS position, the searchlight cannot be
dimmed. The IR filter shall not be used with a 450 watt
lamp installed. The dimming feature of the controllable
searchlight provides a variable light level from 250 to 0
watts to the pilot and copilot through a switch on each
collective grip marked SRCH LT ON - OFF. Push ON OFF BRT DIM to control power to the light and the DIM/
BRT mode selector. When the light is on, the BRT DIM
switch may be moved to select the desired light level. When
the desired level is reached, the switch is released to the
center position. Power to light and control the searchlight is
provided from the dc essential bus through circuit breakers,
marked LIGHTS, CONTR PWR and SRCH CONTR.
The IR filter may be removed for unaided night flight.
2.71.2 Landing Light. One 600-watt landing light is
mounted on the left side beneath the nose section and is
controlled from both collective pitch stick grips (Figure
2-14). The light can be extended 107° from the stowed
position. A dual function switch is used to operate the light.
The LDG LT PUSH ON-OFF switch controls lighting and
EXT, RETR controls light position. When the light is ON
(LDG LT ON advisory light should be on) and the switch
is at EXT detent, the light can be positioned at any point
between stowed and fully extended, or it will continue to
extend until reaching its limit and power is removed. When
the switch is held at RETR the light retracts to the stowed
position. When the light reaches its stowed position, power
is automatically removed from the motor. The LDG LT
PUSH ON-OFF switch must be pushed OFF (LDG LT
ON advisory light should go off). Refer to Chapter 5 for
extend/retract limitations. During extension, the travel
speed is about 12° per second, and during retract, about 30°
per second. Power to light and control the landing light is
supplied from the No. 1 dc primary bus through circuit
breakers, marked LIGHTS, RETR LDG, CONT and
PWR.
2.71.3 Anticollision Lights. This light system contains
four strobes in two separate units, one beneath the aft fuselage and one on top of the aft pylon section. The lights
are controlled by two switches on the upper console (Figure

Change 3

2-75

TM 1-1520-237-10

2-7) labeled ANTI COLLISION LIGHTS UPPER,
BOTH, LOWER and DAY, OFF, NIGHT. The system
consists of a dual power supply and two interchangeable
day/night anticollision lights. The dual supply system provides separate outputs for the aft fuselage light and the
pylon mounted light. Each anticollision light assembly contains two lamps, the upper lamp within a red lens for night
operation and the lower within a clear lens for day operation. Proper operation is selected by placing the switch to
DAY or NIGHT. The desired strobe(s) is selected by placing the switch to UPPER, LOWER or BOTH. If at
BOTH, the lower fuselage and the aft pylon lights will
alternately flash. If the selector switch is placed to UPPER
or LOWER, only that light will flash. To discontinue operation of the anticollision light(s), the DAY-NIGHT
switch is placed to OFF. Power to operate the anticollision
light system is provided from the No. 2 ac primary bus
through a circuit breaker, marked LIGHTS, ANTI COLL.
2.71.4 Position Lights. Position lights (Figure 2-1) are
outboard of the left and right landing gear support and top
tail pylon. The lights are red on the left, green on the right,
and white on the tail. Control of the position lights is
through the upper console panel containing two switches,
marked POSITION LIGHTS, DIM, OFF, BRT, and
STEADY, FLASH. When the intensity switch is placed to
DIM or BRT, all three lights go on at once. If the
STEADY-FLASH switch is placed to FLASH, the three
lights will flash. The STEADY position causes the lights to
remain on continuously. Power to operate the position lights
is provided by No. 2 dc primary bus through a circuit

2-76

Change 3

breaker, marked POS LTS. Infrared position lights are installed within close proximity of the standard position
lights. NVG operation is selected through a toggle switch
on the upper console (Figure 2-7) marked NAV LTS, with
switch positions NORM and IR. Position lights are to be
selected through a switch marked POSITION LIGHTS,
DIM, OFF, or BRT, and mode of operation through a
switch marked STEADY or FLASH. Power for control of
the IR lights is from the No. 2 dc primary bus through a
circuit breaker marked IR LTS.
2.71.5 Formation Lights. These lights (Figure 2-1) are
on top of the main pylon cowling, tail drive shaft cover,
and horizontal stabilator. The system consists of four green
electroluminescent lights. The lights are controlled by a
single rotary selector switch, marked FORMATION LT,
with marked positions OFF and 1 through 5. Position 5 is
the brightest. When NVG operations are required, IR lights
may be used to enhance viewing outside the helicopter. IR
lights are selected through a toggle switch on the upper
console (Figure 2-7) marked, NAV LTS, NORM, and IR.
This switch shares operation with the IR position lights
when operating in a NVG environment. Dimming of the IR
lights is done with the FORMATION LT control, as used
with the electroluminescent formation lights. Selection of
position 1 through 4 causes the IR formation lights to illuminate at the same intensity. Position 5 causes the lights to
illuminate brighter. Power to operate the formation lights is
provided from the No. 2 ac primary bus through two circuit
breakers, marked LIGHTS, FORM LV and HV.

TM 1-1520-237-10

Section XIV FLIGHT INSTRUMENTS
2.72 PITOT-STATIC SYSTEM.
Two electrically-heated pitot tubes with static ports are
aft and above the pilot’s and copilot’s cockpit doors. The
right pitot tube is connected to the pilot’s instruments and
the left pitot tube is connected to the copilot’s instruments.
Tubing connects the pitot tube static pressure ports to the
airspeed indicators and the altimeters. In addition to standard instrumentation, airspeed data is sensed for operation
of stabilator, flight path stabilization, and command instrument system. Refer to Section IX for pitot tube heater system.
2.73 ATTITUDE INDICATING SYSTEM.
Helicopter pitch and roll attitudes are sensed by the pilot’s and copilot’s vertical displacement gyroscopes, that
apply attitude signals to the vertical situation indicators
(VSI) for visual display (Figure 2-9). Signals are applied
through the VERT GYRO select switches to the remote
indicator on the vertical situation indicators. Helicopter
pitch and roll attitudes are shown on the pilot’s and copilot’s vertical situation indicators. The indicator face contains a fixed bar, representing the helicopter, a movable
sphere with a white horizon line dividing the two colors,
white above and black below, a fixed bank angle scale and
a bank index on the moving sphere. Relative position of the
fixed bar (helicopter) and the horizon line indicates the helicopter’s attitude referenced to the earth horizon. A ROLL
trim knob on the lower left of the VSI permits adjustment
of the roll index about 14°66° right and left from zero. A
PITCH trim knob on the lower right of the VSI permits
adjustment of the indicator sphere 14°66° for dive and
7°63° for climb from zero index. If a power failure or
unbalance occurs in the pilot’s or copilot’s vertical displacement gyroscope, a gyroscope power failure flag will
appear, indicating ATT, warning the pilot or copilot that
pitch and roll attitude signals are not being sent to his indicator. To restore attitude information to the indicator, the
pilot or copilot should press his VERT GYRO select
switch on the MODE SEL panel so that ALTR appears in
the switch window. This causes the ATT flag on the indicator to disappear, and pitch and roll signals are supplied
from the operating gyro, restoring attitude information display. Refer to Chapter 3 for description of VERT GYRO
select switch.
2.74 TURN RATE INDICATING SYSTEM.
A 4-minute turn rate (turn and slip) indicator is at the
bottom center of each VSI (Figure 2-9). The pilot’s and
copilot’s indicators operate independently of each other

through TURN RATE switches on the MODE SEL panels. Each system consists of a rate gyro, a turn slip indicator
and a select switch. The VSI contains a moving turn rate
needle and a fixed turn rate scale for indicating rate and
direction of turn. During straight flight the needle is positioned at the center of the scale. When the helicopter turns,
the rate-of-turn signal from the rate gyroscope deflects the
needle in the proper direction to indicate the turn. Amount
of deflection is proportional to the rate-of-turn. A oneneedle width deflection represents a turn of 1.5° per second.
The VSI also contains a slip indicator that shows uncoordinated turns. If a power failure or unbalance occurs in the
pilot’s or copilot’s rate gyroscope, the associated VSI signal will be lost. To restore rate-of-turn information to the
indicator, the pilot or copilot will press the TURN RATE
switch on his MODE SEL panel so that ALTR appears in
the switch window. This applies alternate rate gyroscope
signals from the operating gyroscope to the indicator.
Power to operate the pilot’s turn rate system is provided
from the dc essential bus through a circuit breaker, marked
PILOT TURN DETR. The copilot’s system is powered
from the No. 1 dc primary bus through a circuit breaker,
marked CPLT TURN RATE GYRO. Refer to Chapter 3
for a description of the TURN RATE select switch.
2.75 AIRSPEED INDICATOR.
Two airspeed indicators (Figure 2-9), are installed on the
instrument panel, one each for the pilot and copilot. The
indicators are differential pressure instruments, measuring
the difference between impact pressure and static pressure.
Instrument range markings and limitations are contained in
Chapter 5, Section II, System Limits.
2.76 ALTIMETER/ENCODER AAU-32A.
Two altimeters are installed on the instrument panel
(Figures 2-9 and 2-22). The altimeter encoder functions as
a barometric altimeter for the pilot and a barometric altitude sensor for the AN/APX-100 transponder in mode C.
The copilot’s functions only as a barometric altimeter. The
system is equipped with a continuously operating vibrator
to improve altitude measuring accuracy. The altimeter’s operating range is from -1000 feet to 50,000 feet. The face of
the instrument has a marked scale from zero to nine in
50-foot units. The operating indicators and controls are a
100-foot pointer, 100-foot drum, 1,000-foot drum, 10,000foot drum, barometric pressure set knob, barometric pressure scale window and warning flag. The warning flag is
only used in conjunction with the encoder. A counter window next to the sweep hand contains the three digital drums

Change 9

2-77

TM 1-1520-237-10

2.79 FREE-AIR TEMPERATURE (FAT) INDICATOR.
ENCODER WARNING
FLAG INDICATOR

1000−FOOT
ALTITUDE
COUNTER

CODE
OFF

9 0 1
ALT

1000 FT

100−FOOT
ALTITUDE
COUNTER

BAROMETRIC
SCALE
SET KNOB

The free-air temperature indicator is a direct reading instrument marked FREE AIR, and reads in degrees Celsius.
One FAT indicator is installed through the center windshield on helicopters without center windshield anti-ice
system. On helicopters with center windshield anti-ice system, two indicators are installed through the overhead windows (Figure 2-4).

100−FOOT
ALTITUDE
NEEDLE

100 FT

2
3
2 9 9 0
5 4

2.80 CLOCK.
a. Two clocks (Figure 2-9) are installed on the instrument panel. The elapsed time knob is on the upper right
corner of the clock. The clock is wound and set with a knob
on the lower left corner.

IN. HG .

BAROMETER
PRESSURE
SCALE
AA0520
SA

Figure 2-22. Altimeter Encoder AAU-32A
that rotate to indicate the altitude of the helicopter. Another
window in the upper left section of the instrument face
indicates the normal code operation. When the system fails
to transmit signals to the transponder, a flag marked CODE
OFF will appear in the window. A window on the lower
right section of the instrument face indicates barometric
pressure setting. The barometric pressure set knob is on the
lower left corner of the indicator bezel. Power to operate
the encoder system is provided by the No. 2 dc primary bus
through a circuit breaker, marked PILOT ALTM.
2.77 VERTICAL SPEED INDICATOR.
Two indicators are installed, one each in front of the
pilot and copilot (Figure 2-9), to indicate rate of climb or
descent.
2.78 STANDBY MAGNETIC COMPASS.
A magnetic compass is installed above the instrument
panel on the right center windshield frame (Figure 2-4).
The compass is used as a standby instrument for heading
references. A compass correction card with deviation errors
is installed on the right side of the upper console.

2-78

Change 6

b. Two digital clocks may be installed on the instrument
panel. The clock incorporates a six digit liquid crystal display, 24 hour numerals and sweep second indication. A
battery allows continuous operation for a minimum of one
year when aircraft 28 vdc power is not applied. The clock
has two modes of operation, clock mode (C) and the
elapsed time mode (ET). Power to operate the clock is provided by the No. 1 dc and No. 2 dc primary buses through
circuit breakers marked CPLT ALTM and PILOT ALTM
respectively.
2.81 MASTER WARNING SYSTEM.
Two master caution lights (Figures 2-9 and 2-23) one
each side for the pilot and copilot, marked MASTER
CAUTION PRESS TO RESET, are on the master warning panel. They light whenever a caution light goes on.
These lights alert the pilots and direct attention to the
caution/advisory panel. The master caution lights should be
reset at once to provide a similar indication if a second
condition or malfunction occurs while the first is still
present. The master caution light can be reset from either
pilot position. Four amber warning lights, also on the master warning panel, require immediate action if they go on.
The markings are #1 ENG OUT, #2 ENG OUT, FIRE,
and LOW ROTOR RPM. The LOW ROTOR RPM
warning light will flash at a rate of three to five flashes per
second if rotor rpm drops below 96% RPM R. In addition,
if % RPM R drops below 96% or Ng drops below 55%, a
low steady tone is provided. The low rotor rpm tone is
inhibited on the ground through the left landing gear
weight-on-wheels switch. The engine Ng steady tone is not
inhibited. The ENG OUT warning lights and tone will go
on at 55% Ng SPEED and below. Refer to paragraph 2.14.1

TM 1-1520-237-10

#2 ENG
OUT

#1 ENG
OUT

FIRE

MASTER CAUTION
PRESS TO RESET

LOW ROTOR
RPM

AA0406
SA

Figure 2-23. Master Warning Panel
for description of the FIRE warning lights. Power for the
master caution lights is provided from the No. 1 dc primary
bus through a circuit breaker, marked LIGHTS CAUT/
ADVSY.
2.81.1 Caution/Advisory Light System. The caution/
advisory panel, (Figures 2-9 and 2-24) is on the left of
center of the instrument panel. The caution section (upper
two-thirds) of the panel, indicates certain malfunctions or
unsafe conditions with amber lights. The advisory section
(lower one-third) of the panel shows certain noncritical
conditions with green lights. Each light has its own operating circuit and will remain on as long as the condition that
caused it to light up exists. The caution and advisory lights
are powered by the No. 1 dc primary bus through a circuit
breaker, marked LIGHTS CAUT/ADVSY. Refer to major
systems for a complete description of caution-advisory
panel capsules. Refer to Table 2-3 for a brief description of
each fault.

DIM and TEST, on the lower left of the caution/advisory
panel (Figure 2-24). Placing the switch to TEST simultaneously checks all lights on the caution/advisory and the
master warning panels and #1 and #2 FUEL LOW caution
lights and LOW ROTOR RPM warning lights will flash.
When the pilot’s PILOT FLT rotary intensity control is
moved from the OFF position, placing the BRT/DIMTEST switch to BRT/DIM causes the caution/advisory
lights and master warning lights to change intensity. When
the lights are dim and power is removed, the light intensity
will return to bright when power is reapplied. The TEST
switch position receives power from the No. 1 dc primary
bus through a circuit breaker, marked LIGHTS CAUT/
ADVSY. The BRT/DIM switch position receives power
from the dc essential bus through a circuit breaker, marked
CAUT/ADVSY PNL, on the No. 1 circuit breaker panel.
Dimming of the cockpit indicator lights operates with the
CAUTION panel dimming system.

2.81.2 Caution/Advisory BRT/DIM - TEST Switch.
Testing of the caution/advisory lights is done through a
momentary spring-loaded to center switch marked BRT/
Table 2-3. Caution/Advisory and Warning Light Lighting Parameters
LEGEND

ILLUMINATING PARAMETER OR FAULT
CAUTION CAPSULES

#1 FUEL LOW

Flashes when left fuel tank level is about 172 pounds.

#1 FUEL PRESS

Left engine fuel pressure between engine-driven low-pressure fuel pump and highpressure fuel pump is low.

Change 5

2-79

TM 1-1520-237-10

C
A
U
T
I
O
N

A
D
V
I
S
O
R
Y

#1 FUEL LOW

#1 GEN

#2 GEN

#2 FUEL LOW

#1 FUEL
PRESS

#1 GEN BRG

#2 GEN BRG

#2 FUEL
PRESS

#1 ENGINE
OIL PRESS

#1 CONV

#2 CONV

#2 ENGINE
OIL PRESS

#1 ENGINE
OIL TEMP

AC ESS
BUS OFF

DC ESS
BUS OFF

#2 ENGINE
OIL TEMP

CHIP
#1 ENGINE

BATT LOW
CHARGE

BATTERY
FAULT

CHIP
#2 ENGINE

#1 FUEL
FLTR BYPASS

GUST
LOCK

PITCH BIAS
FAIL

#2 FUEL
FLTR BYPASS

#1 ENGINE
STARTER

#1 OIL
FLTR BYPASS

#2 OIL
FLTR BYPASS

#2 ENGINE
STARTER

#1 PRI
SERVO PRESS

#1 HYD
PUMP

#2 HYD
PUMP

#2 PRI
SERVO PRESS

TAIL ROTOR
QUADRANT

IRCM
INOP

AUX FUEL

#1 TAIL RTR
SERVO

MAIN XMSN
OIL TEMP

INT XMSN
OIL TEMP

TAIL XMSN
OIL TEMP

APU OIL
TEMP HI

BOOST SERVO
OFF

STABILATOR

SAS OFF

TRIM FAIL

LFT PITOT
HEAT

FLT PATH
STAB

IFF

RT PITOT
HEAT

CHIP INPUT
MDL − LH

CHIP
INT XMSN

CHIP
TAIL XMSN

CHIP INPUT
MDL − RH

CHIP ACCESS
MDL − LH

CHIP MAIN
MDL SUMP

APU
FAIL

CHIP ACCESS
MDL − RH

MR DE−ICE
FAIL

MR DE−ICE
FAULT

TR DE−ICE
FAIL

ICE
DETECTED

MAIN XMSN
OIL PRESS

#1 RSVR
LOW

#2 RSVR
LOW

BACK−UP
RSVR LOW

#1 ENG
ANTI−ICE ON

#1 ENG INLET
ANTI−ICE ON

#2 ENG INLET
ANTI−ICE ON

#2 ENG
ANTI−ICE ON

APU ON

APU GEN ON

PRIME BOOST
PUMP ON

BACK−UP
PUMP ON

APU ACCUM
LOW

SEARCH LT
ON

LDG LT ON

#2 TAIL RTR
SERVO ON

CARGO
HOOK OPEN

HOOK ARMED

GPS POS
ALERT

PARKING
BRAKE ON

EXT PWR
CONNECTED

BRT /
DIM

GPS

TEST

AA0354_1B
SA

Figure 2-24. Caution/Advisory Panel (Sheet 1 of 2)

2-80

Change 1

TM 1-1520-237-10

C
A
U
T
I
O
N

A
D
V
I
S
O
R
Y

#1 FUEL LOW

#1 GEN

#2 GEN

#2 FUEL LOW

#1 FUEL
PRESS

#1 GEN BRG

#2 GEN BRG

#2 FUEL
PRESS

#1 ENGINE
OIL PRESS

#1 CONV

#2 CONV

#2 ENGINE
OIL PRESS

#1 ENGINE
OIL TEMP

AC ESS
BUS OFF

DC ESS
BUS OFF

#2 ENGINE
OIL TEMP

CHIP
#1 ENGINE

BATT LOW
CHARGE

BATTERY
FAULT

CHIP
#2 ENGINE

#1 FUEL
FLTR BYPASS

GUST
LOCK

ANTENNA
EXTENDED

#2 FUEL
FLTR BYPASS

#1 ENGINE
STARTER

#1 OIL
FLTR BYPASS

#2 OIL
FLTR BYPASS

#2 ENGINE
STARTER

#1 PRI
SERVO PRESS

#1 HYD
PUMP

#2 HYD
PUMP

#2 PRI
SERVO PRESS

TAIL ROTOR
QUADRANT

ASE

MAIN XMSN
OIL TEMP

INT XMSN
OIL TEMP

TAIL XMSN
OIL TEMP

APU OIL
TEMP HI

BOOST SERVO
OFF

STABILATOR

SAS OFF

TRIM FAIL

LFT PITOT
HEAT

FLT PATH
STAB

IFF

RT PITOT
HEAT

CHIP INPUT
MDL − LH

CHIP
INT XMSN

CHIP
TAIL XMSN

CHIP INPUT
MDL − RH

CHIP ACCESS
MDL − LH

CHIP MAIN
MDL SUMP

APU
FAIL

CHIP ACCESS
MDL − RH

MR DE−ICE
FAIL

MR DE−ICE
FAULT

TR DE−ICE
FAIL

ICE
DETECTED

MAIN XMSN
OIL PRESS

#1 RSVR
LOW

#2 RSVR
LOW

BACK−UP
RSVR LOW

#1 ENG
ANTI−ICE ON

#1 ENG INLET
ANTI−ICE ON

#2 ENG INLET
ANTI−ICE ON

#2 ENG
ANTI−ICE ON

APU ON

APU GEN ON

PRIME BOOST
PUMP ON

BACK−UP
PUMP ON

APU ACCUM
LOW

SEARCH LT
ON

LDG LT ON

#2 TAIL RTR
SERVO ON

AIR COND
ON

CABIN HEAT
ON

ANTENNA
RETRACTED

PARKING
BRAKE ON

EXT PWR
CONNECTED

BRT /
DIM

AUX FUEL

#1 TAIL RTR
SERVO

TEST

AA0354_2
SA

Figure 2-24. Caution/Advisory Panel (Sheet 2 of 2)

2-81

TM 1-1520-237-10

Table 2-3. Caution/Advisory and Warning Light Lighting Parameters (Cont)
LEGEND

ILLUMINATING PARAMETER OR FAULT

#1 ENGINE OIL PRESS

Left engine oil pressure is too low for continued operation.

#1 ENGINE OIL TEMP

Left engine oil temperature is above 150°C.

CHIP #1 ENGINE

Left engine chip detector in scavenge oil system has metal chip or particle buildup.

#1 FUEL FLTR BYPASS

Left engine fuel filter has excessive pressure differential across filter.

#1 ENGINE STARTER

Left engine start circuit is actuated.

#1 PRI SERVO PRESS

First stage pressure is shut off, or has dropped below minimum, or servo pilot valve is
jammed.

TAIL ROTOR
QUADRANT

Goes on when a tail rotor cable is broken or disconnected.

MAIN XMSN OIL TEMP

Main transmission oil temperature is above 120°C.

BOOST SERVO OFF

Indicates loss of second stage hydraulic pressure to the boost servo, or a boost servo jam.

LFT PITOT HEAT

Indicates left pitot heater element is not receiving power with PITOT HEAT switch ON.

CHIP INPUT MDL-LH

Indicates a metal particle has been detected by the chip detector.

CHIP ACCESS MDL-LH

Indicates a metal particle has been detected by the chip detector.

MR DE-ICE FAIL

Indicates a short or open in the main rotor deice system, which will disable the system.

MAIN XMSN OIL PRESS

Main transmission oil pressure is below about 14 psi.

#1 GEN

Left generator is not supplying power to the buses.

#1 GEN BRG

Generator main bearing is worn or has failed.

#1 CONV

Left converter (ac to dc current) has no output.

AC ESS BUS OFF

Indicates that no power (115 vac, phase B) is being supplied to the ac essential bus.

BATT LOW CHARGE

SLAB - Indicates that the voltage on the battery utility bus is at or below 23 vdc. NICAD
- Indicates that the battery charge state is at or below about 40% of full charge state.

GUST LOCK

Indicates the gust lock is not fully disengaged.

#1 OIL FLTR BYPASS

Left engine oil filter pressure differential is excessive.

#1 HYD PUMP

Left hydraulic pump output pressure below minimum.

IRCM INOP

Indicates a malfunction has been detected by the infrared countermeasure system or
infrared countermeasure system is in a cooldown cycle.

ASE

Indicates the ALQ-156 system is being jammed or the ALQ-136, ALQ-144, ALQ-156, or
ALQ-162 system is degraded.

INT XMSN OIL TEMP

2-82

Change 9

Intermediate gear box oil temperature is excessive.

TM 1-1520-237-10

Table 2-3. Caution/Advisory and Warning Light Lighting Parameters (Cont)
LEGEND

ILLUMINATING PARAMETER OR FAULT

STABILATOR

Stabilator system is turned on but is in the manual mode.

FLT-PATH STAB

Indicates that FPS is inoperative in one or more axis.

CHIP INT XMSN

Indicates a metal particle has been detected by the chip detector.

CHIP MAIN MDL SUMP

Indicates a metal particle has been detected by the chip detector.

MR DE-ICE FAULT

Indicates partial failure of the blade deice system. Uneven shedding of ice can be expected.

#1 RSVR LOW

Hydraulic fluid level has dropped below about 60% of full capacity.

#2 GEN

Right generator is not supplying power to the buses.

#2 GEN BRG

Generator main bearing is worn or has failed.

#2 CONV

Right converter (ac to dc current) has no output.

DC ESS BUS OFF

Indicates that no power is being supplied to the dc essential bus.

BATTERY FAULT

Indicates that the battery has exceeded safe operating temperature (overtemperature), or a
battery cell dissimilarity exists. (On helicopters prior to serial number 97-26744)

PITCH BIAS FAIL

(No longer used)

ANTENNA EXTENDED

ECM antenna not fully retracted and at least one of these conditions exist: Helicopter is
below radar altimeter LO bug setting, or power is lost, or AN/APN-209 is turned off or
is removed.

#2 OIL FLTR BYPASS

Right engine oil filter pressure differential is excessive.

#2 HYD PUMP

Right hydraulic pump output pressure below minimum.

ERFS

AUX FUEL

Indicates one or more auxiliary fuel tanks are empty and/or a degraded mode of system
operation.

TAIL XMSN OIL TEMP

Tail gear box oil temperature is excessive.

SAS OFF

Hydraulic pressure supplied to the SAS actuator is below minimum.

IFF

Mode 4 is not capable of responding to interrogation.

CHIP TAIL XMSN

Indicates a metal particle has been detected by the chip detector.

APU FAIL

APU was automatically shut down by the electrical sequence unit.

TR DE-ICE FAIL

Indicates a short or open in a tail rotor blade deice element.

#2 RSVR LOW

Hydraulic fluid level has dropped below about 60% of full capacity.

#2 FUEL LOW

Flashes when right fuel level is about 172 pounds.

#2 FUEL PRESS

Right engine fuel pressure between engine-driven low-pressure fuel pump and highpressure fuel pump is low.

Change 9

2-83

TM 1-1520-237-10

Table 2-3. Caution/Advisory and Warning Light Lighting Parameters (Cont)
LEGEND

ILLUMINATING PARAMETER OR FAULT

#2 ENGINE OIL PRESS

Right engine oil pressure is too low for continued operation.

#2 ENGINE OIL TEMP

Right engine oil temperature is above 150°C.

CHIP #2 ENGINE

Right engine chip detector in scavenge oil system has metal chip or particle buildup.

#2 FUEL FLTR BYPASS

Right engine fuel filter has excessive pressure differential across filter.

#2 ENGINE STARTER

Right engine start circuit is actuated.

#2 PRI SERVO PRESS

Second stage pressure is shut off, or has dropped below minimum, or servo pilot valve is
jammed.

#1 TAIL RTR SERVO

Pressure to the first stage tail rotor servo is below minimum, or servo pilot valve is
jammed.

APU OIL TEMP HI

APU oil temperature is above the maximum.

TRIM FAIL

Indicates that yaw, roll, or pitch trim actuators are not responding accurately to computer
signals.

RT PITOT HEAT

Indicates right pitot heat element is not receiving power with PITOT HEAT switch ON.

CHIP INPUT MDL-RH

Indicates a metal particle has been detected by the chip detector.

CHIP ACCESS MDL-RH

Indicates a metal particle has been detected by the chip detector.

ICE DETECTED

Indicates that ice has been detected.

BACK-UP RSVR LOW

Hydraulic fluid level has dropped below about 60% of full capacity.
ADVISORY CAPSULES

#1 ENG ANTI-ICE ON

Indicates that No. 1 engine anti-ice/start bleed valve is open.

APU ON

APU is operative.

APU ACCUM LOW

APU accumulator pressure is low.

#1 ENG INLET ANTI-ICE
ON

Indicates that No. 1 engine inlet anti-icing air temperature is 93°C or above.

APU GEN ON

APU generator output is accepted and being supplied to the helicopter.

SEARCH LT ON

Either pilot or copilot has selected SRCH LT switch on.

CARGO HOOK OPEN

Indicates that cargo hook load beam is not latched.

AIR COND ON

Power is applied to air conditioner compressor.

PARKING BRAKE ON

Indicates that PARKING BRAKE handle is pulled.

#2 ENG INLET ANTI-ICE
ON

Indicates that No. 2 engine inlet anti-icing air temperature is 93°C or above.

2-84

Change 10

TM 1-1520-237-10

Table 2-3. Caution/Advisory and Warning Light Lighting Parameters (Cont)
LEGEND

ILLUMINATING PARAMETER OR FAULT

PRIME BOOST PUMP ON

Prime boost pump switch is at PRIME or BOOST.

LDG LT ON

Either pilot or copilot has selected LDG LT ON.

HOOK ARMED

The cargo hook release system is armed.

CABIN HEAT ON

Aux heater system is operating.

EXT PWR CONNECTED

Indicates that external power plug is connected to helicopter’s EXT POWER connector.

#2 ENG ANTI-ICE ON

Indicates that No. 2 engine inlet anti-ice/start bleed valve is open.

BACKUP PUMP ON

Backup pump pressure is being supplied.

#2 TAIL RTR SERVO ON

Pressure to 2nd stage tail rotor servo is above minimum.

ANTENNA RETRACTED

ECM antenna fully retracted.

GPS POS ALERT

Indicates that GPS signals are not reliable.

GPS

MASTER WARNING PANEL
#1 ENG OUT

No. 1 engine Ng SPEED is below 55%.

FIRE

Indicates a fire detector has actuated a fire warning circuit.

MASTER CAUTION
PRESS TO RESET

Indicates a caution light on the caution panel has been actuated by failed system.

#2 ENG OUT

No. 2 engine Ng SPEED is below 55%.

LOW ROTOR RPM

Rotor speed is below about 96% RPM R.

Change 9

2-85

TM 1-1520-237-10

Section XV SERVICING, PARKING, AND MOORING
2.82 SERVICING.

2.84.3 Gravity Refueling.

Servicing information is given by systems or components. Points used in frequent servicing and replenishment
of fuel, oil and hydraulic fluid are shown in Figure 2-25.
Fuel and lubricant specifications and capacities are in Table
2-4. Table 2-5 contains a listing of acceptable commercial
fuel.

1. Ground helicopter to fuel truck or other suitable
ground.

2.83 SERVICE PLATFORMS AND FAIRINGS.

3. Remove fuel filler caps and refuel. Refer to
Table 2-4 for fuel quantities.

Service platforms are a part of the engine cowlings, providing access to the engines. Each service platform is about
46 inches long and 18 inches wide. It is capable of supporting a static weight of 400 pounds on any area without
yielding. The platform is made of composite metal and fiberglass with a honeycomb core. The engine cowling is
opened by releasing a latch on the side and pulling outward
on a locking handle. The cowling is opened outward and
down, providing a standing area at the lower section. When
closed, the cowling lock prevents opening in flight.
2.84 FUEL SYSTEM SERVICING.
a. Both tanks (Figure 2-25) may be serviced simultaneously through pressure refueling or closed circuit refueling. They may be serviced individually by gravity refueling
through refueling ports on the left and right sides of the
helicopter.
b. ERFS The external extended range tanks can only
be serviced by gravity refueling through refueling ports on
the forward top of each tank.

2. Plug hose nozzle ground into the helicopter
grounding jack, marked GROUND HERE,
above refueling ports.

2.84.4 Pressure Refueling.
1. Ground helicopter to fuel truck or other suitable
ground.
2. Ground fuel dispenser nozzle to the helicopter
grounding point marked GROUND HERE,
above refueling ports.

CAUTION

Damage to the fuel system could result if
refueling hose pressure exceeds 55 psi
during pressure refueling or 15 psi during
closed circuit refueling.
3. Connect fuel dispenser nozzle to pressure refueling adapter.
NOTE

2.84.1 Fuel Types. Fuels are classified in Table 2-5.
2.84.2 Use of Fuels. Mixing of fuels in fuel tanks.
When changing from one type of authorized fuel to another, for example JP-4 to JP-5, it is not necessary to drain
the helicopter fuel system before adding the new fuel. Fuels
having the same NATO code number are interchangeable.
Jet fuels conforming to ASTM D-1655 specification may
be used when MIL-T-5624 fuels are not available. This
usually occurs during cross-country flights where helicopters using NATO F-44 (JP-5) are refueled with NATO F-40
(JP-4) or Commercial ASTM Type B fuels. Whenever this
condition occurs, the operating characteristics may change
in that lower operating temperature: slower acceleration,
easier starting, and shorter range may be experienced. The
reverse is true when changing from F-40 (JP-4) fuel to F-44
(JP-5) or Commercial ASTM Type A-1 fuels.
2-86

Change 8

The system is designed to restrict fuel flow
to 300 gpm during pressure refueling at a
nozzle pressure of 55 psi and 110 gpm at a
nozzle pressure of 15 psi during closed circuit refueling.
4. Start fuel flow from fuel dispenser and refuel
helicopter.

CAUTION

If fuel is observed flowing from vent, discontinue refueling and make an entry on
DA Form 2408-13-1.

TM 1-1520-237-10

3
7

4
C

APU OIL FILLER CAP
AND DIPSTICK

A2
IN B

L
EP E

L
VE

FULL

LS
UL E

FILL TO
SPILL PLUG

LLUF

OIL LEVEL
INDICATOR

ADD

A1
B
OIL LEVEL INDICATOR

ADD

DDA

T−62T−40−1 APU
FRONT

NO. 1 NO. 2 AND BACKUP
HYDRAULIC PUMP MODULES
LEVEL INDICATOR

(SAME FOR NO. 1
AND NO. 2 ENGINES)

HYDRAULIC FLUID
COLOR
SPEC MIL−H
FULL LEVEL CAPACITY 65 CU. IN. @ 70
REFILL LEVEL CAPACITY 35 CU. IN. @ 70 F

1. & 2. AUXILIARY POWER UNIT

6
7
RED
(REFILL)

OIL FILL
CAP

ENGINE OIL
LEVEL INDICATOR

GREEN
BLUE
(FULL) (EXPANSION)

NOTE
SOME HELICOPTERS MAY HAVE COLORS AS RED
FOR REFILL, GREEN FOR FULL, AND BLACK FOR
EXPANSION AS VIEWED FROM HELICOPTER RIGHT
SIDE

3. INTERMEDIATE GEAR BOX OIL
LEVEL INDICATOR

4. CLOSED CIRCUIT AND PRESSURE
REFUELING PORTS, NO. 1 (LEFT)
FUEL TANK GRAVITY REFUEL
PORT
5. NO. 1 AND NO. 2 ENGINE
OIL LEVEL INDICATOR
6. NO. 1 HYDRAULIC PUMP MODULE
7. BACKUP HYDRAULIC PUMP
MODULE

A2
GTC−P36−150 APU

NO. 1 (LEFT) FUEL
TANK GRAVITY
REFUEL PORT

NO. 1 ENGINE
OIL LEVEL
INDICATOR

5
C
FR

CLOSED
CIRCUIT
REFUEL
PORT CAP

NO. 1 ENGINE LEFT SIDE

PRESSURE
REFUEL
PORT CAP

B
INTERMEDIATE GEAR BOX
AA0324_1B

(SAME FOR NO. 1 AND NO. 2 ENGINES)

Figure 2-25. Servicing Diagram (Sheet 1 of 3)

Change 7

2-87

TM 1-1520-237-10

11
8

TRANSMISSION OIL
LEVEL DIPSTICK

ENGINE OIL
FILLER CAP

OIL
FILLER
CAP

D
OR

FAR SIDE
VIEW
D

NO. 2 ENGINE RIGHT SIDE
(SAME FOR NO. 1 AND NO. 2 ENGINES)

RESERVOIR
QUANTITY
LOW SWITCH

INDICATOR
HANDLE

HANDPUMP
SELECTOR
VALVE

NO. 2 HYDRAULIC
PUMP MODULE
FLUID LEVEL
INDICATOR
(SAME FOR ALL
PUMP MODULES)

INDICATOR
(SERVICEABLE IF
GOLD BAND EXPOSED)

MAIN ROTOR DAMPER

HYDRAULIC PUMP
MODULE REFILL
HANDPUMP

8. MAIN TRANSMISSION OIL FILLER PORT
AND DIP STICK
9. NO. 2 ENGINE OIL FILLER PORT AND
SIGHT GAGE
10. NO. 2 HYDRAULIC PUMP MODULE AND
PUMP MODULE REFILL HANDPUMP

REFILL WITH ONE
QUART WHEN FLUID
REACHES THIS LEVEL

11. MAIN ROTOR DAMPER CHARGE
INDICATOR
12. NO. 2 (RIGHT) FUEL TANK GRAVITY
REFUEL PORT
13. TAIL ROTOR GEAR BOX LEVEL SIGHT
GAGE

Figure 2-25. Servicing Diagram (Sheet 2 of 3)

2-88

OIL LEVEL
INDICATOR

AA0324_2
SA

TM 1-1520-237-10

NO. 2 TANK

NO. 1 TANK
LOOKING AFT

GUIDE
TUBE

HANDPUMP

13
STOW PUMP
ON GRAVITY
FILL DOOR

F
FRONT
SUMP
DRAIN

FUEL TANK GRAVITY
REFUEL PORT

FUEL SAMPLING

TAIL ROTOR
GEAR BOX
AA0324_3

Figure 2-25. Servicing Diagram (Sheet 3 of 3)
5. Once fuel has reached the desired level, remove
the fuel dispenser nozzle from the refueling
adapter and cap pressure fueling adapter.
2.84.5 Fuel Sampling System. Fuel sampling is done
with a thumb-operated handpump (Figure 2-25) containing
5 feet of plastic tubing. The tubing is placed in a guide tube
inside the fuel tank and is directed to the bottom of the
tank. The handpump is stroked and fuel is drawn from the
tank, with contaminants at the bottom. When sampling is
completed, the tubing is emptied, rolled, and stowed with

the pump on the gravity refueling door. ERFS Fuel sampling of the external extended range fuel system is done by
taking the sample with a fuel sampler tube from the sump
drain located at the bottom aft of each tank.
2.85 EXTERNAL AIR SOURCE/ELECTRICAL REQUIREMENTS.
Refer to Chapter 5 for limitations.

Change 5

2-89

TM 1-1520-237-10

2.86 ENGINE OIL SYSTEM SERVICING.

ently added to the engines, drain the oil and
add MIL-L-7808 or MIL-L-23699 oil.
Flushing the system before refilling is not
required.

CAUTION

The engine oil tank (Figure 2-25) is within the main
frame. When the oil level reaches the ADD mark, oil should
be added to bring the level to the full mark on the sight
gage. Wait at least 20 minutes after engine shutdown before
checking engine oil level. Before adding oil, determine
whether system contains MIL-L-7808 oil or MIL-L-23699
oil. If flights of over 6 hours are made, engine oil level
must be at the full line of sight glass before flight.

The helicopter must be level to get accurate oil tank readings. When the helicopter is parked on a slope, the downslope
engine will read higher oil level than actual, and the upslope engine will read
lower.
NOTE
Do not service the engines with DOD-L85734 oil. If DOD-L-85734 oil is inadvert-

Table 2-4. Fuel and Lubricants, Specifications, and Capacities
SYSTEM

SPECIFICATION

CAPACITY

Fuel

Primary: Grade JP-8 (NATO Code F-34)
(Notes 1 and 5)
Alternate: Grade JP-5 (NATO Code F-44)
(Notes 1and 5)
JP-4 (NATO Code F-40) (Note 5)

Main Tanks usable U. S. Gallons of
fuel are: 360 gravity, 359 pressure,
and 356 closed circuit. External
Tank Gravity Refueling: 230 U. S.
Gallons each tank.

WARNING
Lubricating oils MIL-L-23699, DOD-L85734 and MIL-L-7808 contain materials
hazardous to health. They produce paralysis if swallowed. Prolonged contact
may irritate the skin. Wash hands thoroughly after handling. Oils may burn if
exposed to heat or flames. Use only with
proper ventilation.
Engine oil

MIL-L-23699 (NATO Code 0-156)
MIL-L-7808 (NATO Code 0-148)
(Notes 2, 3 and 7)

7 U. S. Quarts

Auxiliary power unit

MIL-L-23699 (NATO Code 0-156)
MIL-L-7808 (NATO Code 0-148)
(Notes 2, 3 and 7)

3 U. S. Quarts
(T-62T-40-1)
2 U. S. Quarts (GTC-P36-150)

Transmission oil

MIL-L-23699 (NATO Code 0-156)
MIL-L-7808 (NATO Code 0-148)
DOD-L-85734 (Notes 2, 3, 6, and 8)

7 U. S. Gallons

2-90

Change 9

TM 1-1520-237-10

Table 2-4. Fuel and Lubricants, Specifications, and Capacities (Cont)
SYSTEM

SPECIFICATION

CAPACITY

Intermediate gear box oil

MIL-L-23699 (NATO Code 0-156)
MIL-L-7808 (NATO Code 0-148)
DOD-L-85734 (Notes 2, 3, 6, and 8)

2.75 U. S. Pints

Tail gear box oil

MIL-L-23699 (NATO Code 0-156)
MIL-L-7808 (NATO Code 0-148)
DOD-L-85734 (Notes 2, 3, 6, and 8)

2.75 U. S. Pints

First stage hydraulic reservoir

MIL-H-83282
MIL-H-5606 (NATO Code H-515)
(Note 4)

1 U. S. Quart

Second stage hydraulic reservoir

MIL-H-83282
MIL-H-5606 (NATO Code H-515)
(Note 4)

1 U. S. Quart

Backup hydraulic reservoir

MIL-H-83282
MIL-H-5606 (NATO Code H-515)
(Note 4)

1 U. S. Quart

Rescue Hoist

Refer to TM 55-1680-320-23 for servicing.

11.5 U. S. Fluid Ounces

SOURCE

PRIMARY OIL

ALTERNATE OIL

APPROVED COMMERCIAL OILS
NOTE

Commercial oils listed below are approved alternates for
engines and gear boxes except as indicated.
U. S. Military Oil
NATO Code No.

MIL-L-23699 or
0-156

MIL-L-7808
0-148

COMMERCIAL OIL

TYPE II

TYPE I

Castrol Inc.

Castrol 5050
Castrol 5000
Aerojet 5

Castrol 399

Turbo Oil 2380

Turbo Oil 2389
Turbo Oil 2391

Hatco Corp.

HATCO
HATCO
HATCO
HATCO

HATCO 1278
HATCO 1280

Mobil Corp.

Mobil Jet Oil II
Mobil Jet Oil 254

Exxon Co.

DOD-L-85734 (Note 7)

Turbo Oil 25

3211
3611
1639
1680

Change 5

2-91

TM 1-1520-237-10

Table 2-4. Fuel and Lubricants, Specifications, and Capacities (Cont)
SOURCE

PRIMARY OIL

ALTERNATE OIL

Royal Lubricants

Royco 555

Royco
Royco
Royco
Royco

Shell Oil Company

Aeroshell 555

Aeroshell 500
Aeroshell 560

500
560
899
899HC

Royco 808

Aeroshell 308

NOTE

2-92

When starting in ambient temperatures below -34°C (-29°F), do not use
JP-5 or JP-8.

When starting in ambient temperatures of -34°C (-29°F) or below,
lubricating oil MIL-L-7808 must be used. It is not advisable to mix
MIL-L-23699 or DOD-L- 85734 oil with MIL-L-7808 oil.

If the type oil being used is not available, another authorized type oil may
be added. When one type oil is mixed with another, it is not necessary to
drain the system and refill with one type oil. No mixing is allowed for cold
temperature operation. For transmissions and gear boxes, when one type of
oil is mixed with another, it is not necessary to drain the system and refill
with one type oil.

For operation below -34°C (-29°F), MIL-H-5606 (NATO Code H-515)
shall be used. Mixing MIL-H-5606 with MIL-H-83282 degrades fireresistant qualities of MIL-H-83282.

Fuel settling time for jet (JP) fuel is 1 hour per foot depth of fuel. Allow the
fuel to settle for the prescribed period before any samples are taken (about
4 hours for proper settling).

DOD-L-85734 oil is the preferred oil for use in the main transmission,
intermediate gearbox, and tail gearbox, except for cold temperature
operation.

DOD-L-85734 oil should not be used in the engines or the auxiliary power
unit (APU). If DOD-L-85734 oil is inadvertently added to the engines or
APU, the system should be drained and the correct oil added. There is no
need to flush the system.

When changing from MIL-L-7808 or MIL-L-23699 oil to DOD-L-85734
(and vice versa), drain the oil from the system and refill with desired oil.
There is no need to flush the system before refilling.

Change 3

TM 1-1520-237-10

Table 2-5. Approved Fuels
SOURCE

PRIMARY/STANDARD
FUEL

U. S. Military Fuel

JP-8

JP-5

JP-4

NATO Code No.

F-34

F-44

F-40

Commercial Fuel
(ASTM-D-1655)

JET A-1

JET A

JET B

American Type A

American JP-4
Arcojet B

American Oil Co.

ALTERNATE FUELS

Atlantic Richfield

Arcojet A-1

Arcojet A

Richfield Div.

Richfield A-1

Richfield A

B. P. Trading

B. P. A. T. K.

B. P. A. T. G.

Caltex Petroleum Corp.

Caltex Jet A-1

Caltex Jet B

City Service Co.

CITCO A

Continental Oil Co.

Conoco Jet-60

Conoco Jet-50

Conoco JP-4

Exxon Co. U. S. A.

Exxon A-1

Exxon A

Exxon Turbo Fuel B

Gulf Oil

Gulf Jet A-1

Gulf Jet A

Gulf Jet B

Mobil Oil

Mobil Jet A-1

Mobil Jet A

Mobil Jet B

Philjet A-50

Philjet JP-4
Aeroshell JP-4

Phillips Petroleum
Shell Oil

Aeroshell 650

Aeroshell 640

Sinclair

Superjet A-1

Superjet A

Standard Oil Co.

Jet A-1 Kerosene

Jet A Kerosene

Chevron

Chevron A-1

Chevron A-50

Chevron B

Texaco

Avjet A-1

Avjet A

Texaco Avjet B

76 Turbine Fuel

Union JP-4

NATO F-44

NATO F-40

Union Oil
International Fuel

NATO F-34

Belgium
Canada

BA-PF-2B
3-6P-24C

3GP-22F

Denmark

JP-4 MIL-T-5624

France

Air 3407A

Germany

UTL-9130-007
UTL-9130-010

Greece

VTL-9130-006
JP-4 MIL-T-5624

Italy

AMC-143

AA-M-C-1421

Netherlands

D. Eng RD 2493

JP-4 MIL-T-5624

Norway

JP-4 MIL-T-5624

Portugal

JP-4 MIL-T-5624

Turkey

JP-4 MIL-T-5624

Change 10

2-93

TM 1-1520-237-10

Table 2-5. Approved Fuels (Cont)
SOURCE

PRIMARY/STANDARD
FUEL

United Kingdom
(Britain)

ALTERNATE FUELS
D. Eng RD 2498

D. Eng RD 2454

NOTE

Commercial fuels are commonly made to conform to
American Society for Testing and Materials (ASTM)
Specification D 1655. The ASTM fuel specification does
not contain anti-icing additives unless specified. Icing
inhibitor conforming to MIL-I-85470 or MIL-I-27686
(Commercial name PRIST) shall be added to commercial and NATO fuels, not containing an icing inhibitor,
during refueling operations, regardless of ambient temperatures. Icing inhibitor conforming to MIL-I-85470 is
replacing the MIL-I-27686 version. The use of MIL-I27686 icing inhibitor is acceptable until all supplies are
depleted. Adding PRIST during refueling operation shall
be done using accepted commercial mixing procedures.
The additive provides anti-icing protection and also
functions as a biocide to kill microbial growths in helicopter fuel systems.
2.87 APU OIL SYSTEM SERVICING.
NOTE
Do not service the APU with DOD-L-85734
oil. If DOD-L-85734 oil is inadvertently
added to the APU, drain the oil and add
MIL-L-7808 or MIL-L-23699 oil. Flushing
the system before refilling is not required.
a. The APU oil supply is in the APU gear box assembly. The sump filler/oil dipstick port (T-62T-40-1) or cap
and fill to spill plug (GTC-P36-150) (Figure 2-25) are on
the left side of the gear box housing.
b. When the APU is cool to the touch the COLD side of
the dipstick may be used, if the APU is hot to the touch the
HOT side of the dipstick may be used.
2.88 HANDPUMP RESERVOIR SERVICING.

CAUTION

Do not allow reservoir level to fall below
refill line.
2-94

Change 10

Servicing of the refill handpump is done when fluid level
decreases to the refill line on the fluid level sight gage, on
the side of the pump tank. When fluid level decreases to the
refill line, 1 quart of hydraulic fluid can be poured into the
reservoir after removing the refill cap. Handpump reservoir
level should be replenished only in 1 quart units.
2.89 HYDRAULIC SYSTEMS SERVICING.
Reservoirs (Figure 2-25) for the hydraulic systems are
on the hydraulic pump modules. Fluid level sight gages are
visible on the side of each pump. All hydraulic pump reservoir capacities are 1 U. S. quart to the blue (black on
some pumps) mark. When the indicator reaches the red area
(refill) point, 2/3 of a pint is required to return the indicator
to the green mark. The fluid level indication is the 1/8 inch
wide gold band at the outboard edge of the level piston. To
refill the reservoirs, the fluid is supplied from the manual
handpump. After flight, fluid in hydraulic systems will be
hot. Piston movement of up to 3/8 inch into the blue (black
on some pumps) (overfill) zone is acceptable. When piston
is beyond this limit, bleed off enough fluid to bring piston
back to 3/8 inch above fill limit. To replenish the pump
reservoir fluid, do the following:
1. Unscrew handpump lid and pour in clean hydraulic fluid, MIL-H-83282, until pump is full.

TM 1-1520-237-10

Make sure you can always see oil in pump reservoir window while servicing, so not to pump
air into pump module’s reservoir. Keep filling.
2. Make sure pump cover is clean, then screw lid
on tight.
3. Turn selector valve to desired reservoir to be
filled. OUT 1 is left pump module, OUT 2 is
right pump module, and OUT 3 is backup
pump module.
4. While holding selector valve handle down,
crank pump handle on handpump clockwise
and fill desired hydraulic pump module until
forward end of piston in reservoir window is at
forward end of green decal on reservoir housing.
5. Check that reservoirs stay full (forward end of
piston at forward end of green decal), with fluid
at ambient temperature 1 hour after flight.

NOTE
Remove the dipstick, clean and reinsert to
obtain correct reading.
a. Single scale dipstick is for checking cold oil levels.
Wait at least 2 hours after shutdown to check oil. If oil
level must be checked when hot (immediately to 1/2 hour
after shutdown), oil level will read about 1/2 inch low
(halfway between full and add mark or 1/2 inch below add
mark).
b. Dual scale dipstick is for checking cold or hot oil
levels. Use appropriate scale when checking oil level. Read
hot side of dipstick when checking hot oil (immediately to
1/2 hour after shutdown), or cold side of dipstick when
checking cold oil (at least 2 hours after shutdown).
2.92 TAIL AND INTERMEDIATE GEAR BOX SERVICING.

6. Make sure area remains clean during procedure.

The intermediate gear box oil level sight gage (Figure
2-25) is on the left side of the gear box. The tail gear box
oil level sight gage is on the right side.

7. Stow selector valve handle in OUT 4 (capped
off) position.

2.93 PARKING.

8. Turn on electrical power.
9. Check caution panel for #1 RSVR LOW, #2
RSVR LOW, and BACK-UP RSVR LOW
lights are off.
2.90 RESCUE
SERVICING.

HOIST

LUBRICATION

SYSTEM

Servicing of the rescue hoist lubrication system consists
of replacing automatic transmission fluid in the boom head
and the gear box (Figure 4-25) until oil level sight gages
indicate full.
2.91 MAIN TRANSMISSION OIL SYSTEM SERVICING.
The transmission oil supply is in the sump case with the
filler port and dipstick gage (Figure 2-25), on the right rear
of the main module. When filling is required, oil is poured
through the filler tube on the main module case, and oil
level is checked by a dipstick, marked FULL and ADD, or
FULL COLD and ADD on one side of the dipstick and
FULL HOT and ADD on the other side. Check oil level as
follows:

The methods used to secure the helicopter for temporary
periods of time will vary with the local commands. The
minimum requirements for parking are: gust lock engaged
and wheel brakes set, tailwheel locked, and wheels properly
chocked. For extended periods of time, engine inlet covers,
exhaust covers, and pitot covers should be installed, and
stabilator slewed to 0°. When required, the ignition system
and the doors and window should be locked.
2.94 PROTECTIVE COVERS AND PLUGS.
The covers and plugs (Figure 2-26) protect vital areas
from grit, snow, and water. The protected areas are avionics
compartment air inlet, engine air inlet/accessory bay, engine and APU exhausts, pitot tubes, IRCM transmitter and
APU air inlet and main transmission oil cooler exhaust.
Covers and plugs should be installed whenever the helicopter is to be on the ground for an extended period of time.
Each cover may be installed independently of the others.
2.95 MOORING.
Mooring fittings are installed at four points on the helicopter (Figure 2-26). Two fittings are at the front of the
fuselage, one above each main landing gear strut, and two

Change 9

2-95

TM 1-1520-237-10

at the rear, one attached to each side of the aft transition
section. These fittings are used to tie down the helicopter
when parked, and wind conditions require it.

3. Repeat steps 1. and 2. for each remaining blade.
4. Turn blade to about 45° angle to centerline of
helicopter and engage gust lock.

2.95.1 Mooring Instructions. Refer to TM 1-1500250-23 for mooring instructions.
CAUTION

2.95.2 Main Rotor Tiedown. Tiedown of the main rotor should be done when the helicopter will be parked for a
period of time or when actual or projected wind conditions
are 45 knots and above. To tiedown main rotor blades, do
this:
1. Turn rotor head and position a blade over centerline of helicopter. Install tiedown fitting into
receiver while pulling down on lock release
cable. Release cable when fitting is installed in
blade receiver.
2. Uncoil tiedown rope.

2-96

Change 1

Do not deflect main rotor blade tips more
than 6 inches below normal droop position when attaching tiedowns. Do not tie
down below normal droop position.
5. Attach tiedown ropes to helicopter as shown in
Figure 2-26. To release tiedown fitting, pull
down on lock release cable and remove fitting
from blade.

TM 1-1520-237-10

PITOT TUBE COVER
AND WARNING STREAMER
(LEFT AND RIGHT SIDE)

APU EXHAUST
PLUG

IRCM TRANSMITTER
APU AIR INLET,
AND TRANSMISSION
OIL COOLER COVER

ENGINE EXHAUST
PLUGS (LEFT AND
RIGHT SIDE)

ROTOR BLADE
TIEDOWN ROPE
(ON EACH BLADE)

HELICOPTER TAIL
TIEDOWN CABLE
(LEFT AND RIGHT
SIDE)

ENGINE AIR INLET / ACCESSORY
BAY COVER
(LEFT AND RIGHT SIDE)

AVIONICS
COMPARTMENT
AIR INLET COVER

B
HELICOPTER TIEDOWN CABLE
(LEFT AND RIGHT SIDE)
MAIN ROTOR
BLADE

WARNING
STREAMER

FITTING ASSEMBLY

LOCK RELEASE
CABLE

LANDING GEAR
DRAG STRUT

RECEIVER

LOCK ASSEMBLY
TIEDOWN ROPE

WARNING
STREAMER

LOCK RELEASE
CABLE

TIEDOWN LINE
(LEFT AND
RIGHT SIDE)
AA0522A
SA

Figure 2-26. Mooring

2-97/(2-98 Blank)

TM 1-1520-237-10

CHAPTER 3
AVIONICS
Section I GENERAL
3.2 AVIONICS EQUIPMENT CONFIGURATION.

3.1 DESCRIPTION.
The avionics subsystem consist of the communications
equipment providing VHF-AM, VHF-FM, and UHF-AM
communications. The navigation equipment includes, LFADF, VOR, ILS, marker beacon, Doppler UH , Doppler/
GPS UH , or Integrated Inertial Navigation System
EH VHF-FM homing is provided through the No. 1
VHF-FM communication radio. Transponder equipment
consists of a receiver-transmitter with inputs from barometric altimeter for altitude fixing. Absolute height is provided
by a radar altimeter. Each antenna will be described with its
major end item, and locations as shown in Figure 3-1.

Equipment configuration is as shown in Table 3-1.
3.3 AVIONICS POWER SUPPLY.
Primary power to operate the avionics systems is provided from the No. 1 and No. 2 dc primary buses and the dc
essential bus, and No. 1 and No. 2 ac primary buses (Figure
2-20). When operating any of the avionics equipment, helicopter generator output must be available or external ac
power connected. Function selector switches should be at
OFF before applying helicopter power.

Table 3-1. Communication/Navigation Equipment

FACILITY

NOMENCLATURE

USE

RANGE

CONTROL
LOCATION

Intercommunication system

Interphone
control
C-6533/
ARC

Intercommunication between crewmembers and control of navigation
and communication radio.

Stations
within helicopter

Cockpit
lower
console,
crewchief/
gunner’s stations
and troop commander’s station
at center of cabin
overhead
with
handset

REMARKS

Change 1

3-1

TM 1-1520-237-10

Table 3-1. Communication/Navigation Equipment (Cont)

FACILITY

NOMENCLATURE

USE

FM communications (If installed) UH

Radio set
AN/ARC114A
VHF-FM
No. 1

Two-way voice communications;
FM and continuous-wave homing
frequency range 30 through 75.95
MHz.

FM communications (If installed) UH

Radio Set
AN/ARC114A
VHF-FM
No. 2

Same as No. 1 VHF-FM, except no
homing is provided.

VHF communications (If
installed) UH

Radio Set
AN/ARC115A
VHF-AM

Two-way voice communications in
the frequency range of 116.000
through 149.975 MHz.

*Line
sight

Lower console

Radio Set AN/
ARC-115 may
be installed on
some helicopters.

VHF AM and
FM communications (If installed)

Radio Set
AN/ARC186(V)
VHFAM/FM

Two-way voice communications in
the frequency range 30.0 through
87.975 MHz and 116.0 through
151.975 MHz range. 108.0 through
115.975 MHz receive only.

*Line
sight

Lower console
UH , ECM operator’s station

VHF-FM No.
2 Provisional.

FM communications

Radio Set
AN/ARC201
VHF-FM

Two-way voice communications,
homing, frequency hopping in 30.0
- 87.975 MHz range.

*Line
sight

Improved Frequency Modulation Amplifier

IFM Amplifier AM7189A/
ARC

Variable RF power amplifier; increases output from FM 1 (2.5, 10
or 40 watts out.)

3-2

Change 8

RANGE

*Line
sight

CONTROL
LOCATION
Lower console

REMARKS

FM No. 1
transmitter
may be used
only in helicopters serial
No. 79-23273
and
subsequent
and
those helicopters modified
by MOD 99122 and 99122-1.

Lower console

VHF-FM No.
1/2 and MEP
VHF-FM.

FM 1 ARC-201
control box

Amplifier control C-11188A
used
only
when ARC186 is installed EH .

TM 1-1520-237-10

Table 3-1. Communication/Navigation Equipment (Cont)

FACILITY

NOMENCLATURE

USE

RANGE

UHF communications

RadioTransmitter
Radio, RT1167/ARC164(V)
UHF-AM

Two-way voice communications in
the frequency range of 225.000 to
399.975 MHz.

Radio, RT1167C/
ARC164(V)
UHF-AM

HAVE QUICK

Radio, RT1614/ARC164(V)
UHF-AM

HAVE QUICK II

Tunable
diplexer EH

TD-1336/A

Allows narrow band use of guard
channel.

High
frequency communications

Radio Set
AN/ARC220

Two way voice communications in
the frequency range of 2 to 29.9999
MHz.

*Over the
horizon

Lower console

Voice security
system

TSEC/
KY-58

Secure communications.

Not applicable

Lower console

Can be used
with
FM1,
FM2
and
UHF-AM.

Voice security
system

TSEC/KY100

Secure communications.

Not applicable

Rear lower console

Used with HF
AN/ARC-220.

Automatic direction finding

Direction
Finder Set
AN/
ARN-89 (if
installed)
AN/ARN149(V) (if
installed)

Radio range and broadcast reception; automatic direction finding
and homing in the frequency range
of 100 to 3000 kHz.

*50 to 100
miles range
signals.

Lower console

AN/ARN149(V)
tunable, 100 to
2199.5 kHz.

*Line
sight

CONTROL
LOCATION

REMARKS

Lower console
UH , DF operator’s station EH

Lower console
UH

Beneath seat of
copilot

Change 8

3-3

TM 1-1520-237-10

Table 3-1. Communication/Navigation Equipment (Cont)

FACILITY

NOMENCLATURE

USE

VOR/LOC/
GS/MB
receiving set

Radio Receiving Set
AN/ARN123(V) (if
installed)
AN/ARN147(V) (if
installed)

VHF navigational aid, VHF audio
reception in the frequency range of
108 to 117.95 MHz and marker
beacon receiver operating at 75
MHz.

Doppler navigation set UH

Doppler
Navigation
Set
AN/
ASN-128

Provides present position or destination navigation information in
latitude and longitude (degrees and
minutes) or Universal Transverse
Mercator (UTM) coordinates.

Lower console

Doppler/GPS
navigation set

Doppler/
GPS Navigation Set
AN/ASN128B

Provides present position or destination navigation information in
latitude and longitude (degrees and
minutes) or Military Grid Reference System (MGRS) coordinates.
Combined Mode is prime (default)
mode of operation where the GPS
updates Doppler present position.
When Doppler is in memory (if not
available), the system switches to
GPS only mode. If GPS is jammed
and/or becomes unavailable, the
system automatically switches to
the Doppler only mode. These
modes of operation may also be selected manually.

Lower console

AN/ASN132(V)

Navigational Aid.

Lower console

Magnetic
heading indications

Gyro Magnetic Compass AN/
ASN-43

Navigational Aid.

Lower console

Identification
friend or foe

Transponder Set AN/
APX100(V)

Transmits a specially coded reply
to a ground-based IFF radar Interrogator system.

Integrated Inertial Navigation
System
(IINS)

RANGE

*Line
sight

CONTROL
LOCATION
Lower console

3-4

Change 4

*Line
sight

Lower console

REMARKS

AN/ARN147(V)
tunable, 108 to
126.95 kHz.

Doppler
ONLY or GPS
ONLY navigation is selectable from
CDU.

TM 1-1520-237-10

Table 3-1. Communication/Navigation Equipment (Cont)

FACILITY

Absolute
timeter

al-

NOMENCLATURE
Radar Altimeter AN/
APN-209

USE

Measures absolute altitude.

RANGE

CONTROL
LOCATION

0 to 1500
feet

Instrument panel

REMARKS

NOTE

*Range of transmission or reception depends upon many variables including weather
conditions, time of day, operating frequency, power of transmitter and altitude of the
helicopter.

Change 4

3-5

TM 1-1520-237-10

TROOP
COMMANDERS VHF / FM

UH60A/UH60L HELICOPTERS

VHF−FM
NO. 1 /
VHF−AM

VHF−FM NO. 2
ANTENNA

TRANSPONDER (IFF)
ANTENNA (TOP)

GPS ANTENNA

INFRARED
COUNTERMEASURE
TRANSMITTER

VOR / LOC ANTENNA
(SAME BOTH SIDES)

HF ANTENNA

VHF−FM HOMING ANTENNA
(SAME BOTH SIDES)

RADAR WARNING
ANTENNA
GLIDE
SLOPE
ANTENNA

DOPPLER
ANTENNA

RADAR
WARNING
ANTENNA

MARKER
BEACON
ANTENNA

RADAR
WARNING
ANTENNA

LF / ADF LOOP
ANTENNA
UHF COMM
ANTENNA

RADAR
ALTIMETER
ANTENNA

VHF−FM NO. 2, VHF−AM
LF / ADF SENSE ANTENNA

TRANSPONDER (IFF)
ANTENNA (BOTTOM)

BOTTOM VIEW

RADAR
WARNING
ANTENNA

AA0355_1
SA

Figure 3-1. Antenna Arrangement (Sheet 1 of 2)

3-6

Change 5

TM 1-1520-237-10

MISSION VHF / FM
2ND BITE ANTENNA

EH60A HELICOPTERS

VHF−FM NO. 1
DF ANTENNA

INFRARED
COUNTERMEASURE
TRANSMITTER

DF ANTENNA

VOR / LOC ANTENNA
(SAME BOTH SIDES)

TRANSPONDER (IFF)
ANTENNA (TOP)

VHF−FM HOMING ANTENNA
(SAME BOTH SIDES)

MEP VHF−FM /
UHF VOICE LINK
ANTENNA

RADAR WARNING
ANTENNA

ALQ−156
ANTENNA

GLIDE
SLOPE
ANTENNA

RADAR
WARNING
ANTENNA

MARKER
BEACON
ANTENNA

RADAR
WARNING
ANTENNA

ALQ−156
ANTENNA
DF ANTENNA

LF / ADF LOOP
ANTENNA

ALQ−162
ANTENNA

MEP UHF
DATA LINK
ANTENNA

ECM
ANTENNA

VHF−FM
NO. 2
AIRCRAFT
UHF COMM
ANTENNA
ALQ−156
ANTENNA

RADAR
ALTIMETER
ANTENNA

TAC
ANTENNA
VHF−AM LF / ADF
SENSE ANTENNA

ALQ−156
ANTENNA
TRANSPONDER (IFF)
ANTENNA (BOTTOM)

DF
ANTENNA

RADAR
WARNING
ANTENNA

BOTTOM VIEW

ALQ−162
ANTENNA
AA0355_2
SA

Figure 3-1. Antenna Arrangement (Sheet 2 of 2)

Change 5

3-7

TM 1-1520-237-10

Section II COMMUNICATIONS
3.4 INTERCOMMUNICATION SYSTEM C-6533()
/ARC.
Five intercommunication system (ICS) controls provide
interior intercommunication capability between crew members and with the troop commander’s position. They also
provide a means by which the pilot and copilot may select
and control associated radio equipment for voice transmission and reception. Additional audio circuits may also be
selected for constant monitoring. When the communication
control is operated in conjunction with equipment listed in
Table 3-1, it is used to select associated radio equipment
for voice operations. The operator may select any one of
five transmitters (Figure 3-2), and/or any or all of five receivers to monitor. Four direct-wired audio circuits allow
continuous monitoring. Hands-free intercommunication is
provided by a hot mike feature. An exterior jack is to the
front and below each gunner’s window. When the
walkaround cord is connected to it, the crewchief can communicate with the interior of the helicopter or with the other
exterior jack through thecrewchief/gunner’s control panels.
A placard installed on the instrument panel and above each
troop-cargo compartment ICS station control panel indicates which receiver is selected when a selector switch is
placed ON. Power for the intercommunication system is
provided from the dc essential bus through circuit breakers,
marked ICS PILOT and ICS COPILOT.
3.4.1 Controls and Functions. Controls for the
intercom/radios are on the front panel of the unit (Figure
3-2). The function of each control is as follows:

CONTROL/
INDICATOR
Receiver selector
switches (ON)
1
2
3
4
5
AUX

FUNCTION

RECEIVER SELECTOR
SWITCH

C
O
M
M

AUX

NAV

OFF

VOL

ICS

3
4
5

C
O
N
HOT MIKE T

OFF

TRANSMITTER
SELECTOR
SWITCH

AA0523
SA

Figure 3-2. Intercommunication Control Panel
C-6533/ARC

CONTROL/
INDICATOR
NAV
VOL control
Transmitter selector switch
ICS
1
2

Change 10

2
1

5
HOT MIKE
switch

FUNCTION
Connects ADF/Marker Beacon audio to the headphone.
Adjusts headphone volume level.

Enables intercom operation when
keyed.
Enables FM 1 transmission when
keyed.
Enables UHF transmission when
keyed.
Enables VHF transmission when
keyed.
Enables FM 2 transmission when
keyed (provisions).
Enables HF transmission when
keyed.
Enables intercom transmission
without manual key.

3.4.2 Intercommunication Keying System. Keying
of the ICS system is done by these controls:
a. Pilot or copilot station. An ICS or RADIO switch on
the top of each cyclic stick, or by a switch on the floor at
the pilot’s left and the copilot’s right foot (Figure 2-4).

Pages 3-8.1 through 3-10 deleted.
3-8

OFF

Connects FM 1 receiver to the
headphone.
Connects UHF receiver to the headphone.
Connects VHF receiver to the headphone.
Connects FM 2 receiver to the
headphone.
Connects HF receiver to the headphone.
Connects VOR/LOC audio to the
headphone.

TM 1-1520-237-10

b. Crewchief/Gunner and Left Gunner. A pushbutton at
the end of the ICS cord or the exterior walkaround cord,
and foot switches on each side of the helicopter at the
crewchief/gunner’s and left gunner’s station.
c. Troop commander. A push switch on the handset at
the troop commander’s station.
3.4.3 Modes of Operation.
3.4.3.1 Primary Operation Check. There are several
methods of intercommunication operation. In all cases, no
operator action is required to receive intercom signals other
than adjusting the VOL control for a comfortable level at
the headset.
3.4.3.2 Intercommunication (All Stations).
1. Transmitter selector switch ICS for pilot and
copilot when using foot switch, any position
when using cyclic switch, ICS for crewchief/
gunner, gunner, and troop commander.
2. Key switch - ICS switch on pilot’s or copilot’s
cyclic, or foot switch at pilot’s, copilot’s or
crewchief/gunner, gunner positions, or push-totalk button on crewchief/gunner’s ICS cord,
push-to-talk switch on troop commander handset press, speak into microphone and listen for
sidetone, release to listen.

2. RADIO push-to-talk switch on cyclic stick, or
foot-operated push-to-talk switch - Press; speak
into microphone while holding switch; release
to listen.
3.4.4.1 Crewchief/Gunner.
1. Transmitter selector - Desired position, 1
through 5.
2. Push-to-talk switch on headset-microphone
cord, or foot-operated push-to-talk switch Press, speak into microphone while holding
switch, release to listen.
3.4.4.2 Troop commander.
1. Transmitter selector - Desired position 1
through 5.
2. Transmitter key switch on handset - Press,
speak into microphone while holding switch,
release to listen.
3.4.5 Receiver Selection.
1. Receiver selector switch(es) - ON as desired.
2. Adjust volume to a comfortable listening level.
3.5 DELETED.

3.4.3.3 External Radio Communication. All stations
of the helicopter are capable of external radio communications.

3.6 DELETED.

3.4.4 Pilot and Copilot.

Radio Set AN/ARC-186(V) (Figure 3-5) is a lightweight
multichannel airborne radio communications set, which
provides transmission, reception and retransmission of amplitude modulated (AM), frequency modulated (FM) radio

1. Transmitter selector - Desired position, 1
through 5.

3.7 RADIO SET AN/ARC-186(V).

Change 10

3-11

TM 1-1520-237-10

communications, and FM directional finding (homing) with
installation of other associated equipment. AM reception
only is provided on frequencies between 108.000 and
115.975 MHz. Installation of the AN/ARC-186(V) in the
UH-60A helicopter is a VHF-AM and/or VHF-FM installation(s). The transceiver has a tunable main receiver and
transmitter which operates on any one of 1,469 AM discrete channels, each spaced 25 kHz apart within the frequency range of 116.000 through 151.975 MHz, or 30.000
through 87.975 MHz FM, providing 2,319 channels. FM
homing operations within the 30 through 87.975 MHz band.
The fixed guard channels are between 116.000 and 151.975
MHz AM (usually 121.500 MHz), and between 30.000 and
87.975 MHz FM (usually 40.500 MHz). The guard frequencies are preset and only require selection by the
frequency/emergency select switch. Frequencies can be
preset for 20 channels. VHF-AM installations cannot be
used to transmit VHF-FM signals. If an AM frequency is
selected on an FM only installation, an audible tone would
be heard, warning the pilot of an out-of-band frequency
selection. The same is true in the case of selection of an FM
frequency on an AM installation. Keying the microphone
for voice transmission when in DF (homing) mode will
disable the homing function while the mic is keyed. In DF
mode, audio reception is distorted. When using secured
speech and EMER FM or AM is selected, secure speech
function will be disabled to enable normal voice communications. Power to operate the AN/ARC-186(V) radio is provided from the No. 2 dc primary bus through a circuit
breaker, marked VHF-AM for the AM radio, and from the
No. 1 and No. 2 dc primary buses, respectively, through
circuit breakers marked NO. 1 VHF-FM and NO. 2
VHF-FM for the No. 1 and No. 2 VHF-FM radios.
3.7.1 Antennas.
a. UH The VHF-AM antenna is on top of the tail rotor
pylon (Figure 3-1). The antenna operation is shared with
No. 1 VHF-FM. The No. 2 VHF-FM communications antenna is within the leading edge fairing of the tail pylon
drive shaft cover. The two FM homing antennas used with
the No. 1 VHF-FM radio are on each side of the helicopter
fuselage, just behind the cockpit doors.
b. EH The VHF-AM antenna is under the nose section
(Figure 3-1). The antenna operation is shared with ADF
sense. No. 1 VHF-FM communications antenna is within
the leading edge fairing of the tail pylon drive shaft cover.
No. 2 VHF-FM communications antenna is under the nose,
forward of VHF-AM antenna. The two FM homing antennas used with the No. 1 VHF-FM radio are on each side of
the helicopter fuselage, just behind the cockpit doors.

3-12

3.7.2 Controls and Functions. Controls for the AN/
ARC-186(V) are on the front panel of the unit (Figure 3-5).
The function of each control is as follows:
CONTROL
0.025
selector

FUNCTION
MHz

Rotary switch. Selects rt frequency
in 0.025 MHz steps. Clockwise
rotation increases frequency.

0.1 MHz selector

Rotary switch. Selects rt frequency
in 0.1 MHz steps. Clockwise
rotation increases frequency.

1.0 MHz selector

Rotary switch. Selects rt frequency
in 1.0 MHz steps. Clockwise
rotation increases frequency.

10 MHz selector

Rotary switch. Selects rt frequency
in 10 MHz steps from 30 to 150
MHz. Clockwise rotation increases
frequency.

Preset channel
selector

Selects preset channel from 1 to 20.
Clockwise
rotation
increases
channel number selected.

Volume control

Potentiometer. Clockwise rotation
increases volume.

Squelch disable/
tone select

Three-position switch. Center
position enables squelch, SQ DIS
position
disables
squelch,
momentary
TONE
position
transmits tone of approximately
1000 Hz.

Frequency
control/
emergency select
switch

Four-position rotary switch. PRE
position enables preset channel
selection, MAN position enables
manual
frequency
selection,
EMER AM or FM selects a
prestored guard channel (FM not
used in helicopters with panelmounted transceiver).
NOTE

Selecting EMER AM or FM
automatically disables the secure
speech function and enables normal
voice communication.
Mode select
switch

Three-position rotary switch. OFF
position
disables
receiver/
transmitter, TR position enables
transmit/receive
modes.
DF
position enables FM homing.

TM 1-1520-237-10

10.0 MHZ
INDICATOR

10.0 MHZ
SELECTOR

0.1 MHZ
SELECTOR

1.0 MHZ
INDICATOR

0.1 MHZ
INDICATOR

1.0 MHZ
SELECTOR

0.025 MHZ
INDICATOR

0.025 MHZ
SELECTOR
PRESET CHANNEL
INDICATOR

PRESET CHANNEL
SELECTOR

VOLUME
CONTROL
FREQUENCY
CONTROL /
EMERGENCY
SELECT SWITCH

CHAN

EMER
FM
AM

V
H
F

MAN
PRE
S
Q

D
I
S

20
VOL
SNAP−ON
COVER
T
O
N
E

AM
SQUELCH

MODE
SELECT
SWITCH

SQUELCH
DISABLE /
TONE
SELECT

AM
SQUELCH
CONTROL

LOCK
OUT
AM

OFF

MEM
LOAD

FM
SQUELCH

BANDWITH /
MEM LOAD
SWITCH

1
2
3
4
5

FM SQUELCH
CONTROL

6
7

11
12

8
9

13
14

18
19

BAND LOCKOUT
SWITCH

(PANEL−MOUNTED TRANSCEIVER)

AA0361_1
SA

Figure 3-5. VHF Control AN/ARC-186(V) (Sheet 1 of 2)

3-13

TM 1-1520-237-10

VOLUME
CONTROL

10.0 MHz
INDICATOR

1.0 MHz
SELECTOR

10.0 MHz
SELECTOR

SQUELCH
DISABLE /
TONE SELECT

V
O
L
S
Q
D
I
S
FM

0.1 MHz
INDICATOR

1.0 MHz
INDICATOR

0.025 MHz
INDICATOR

0.1 MHz
SELECTOR

T
O
N
E

C
O
M
M

EMER
AM

PRESET

MAN LOAD
PRE

0.025 MHz
SELECTOR

9
TR

FREQUENCY
CONTROL
SELECT
SWITCH

OFF

LOAD
PUSHBUTTON
SWITCH

MODE
SELECTOR
SWITCH

PRESET
CHANNEL
INDICATOR

PRESET
CHANNEL
SELECTOR
THUMBWHEEL
AA0361_2A
SA

Figure 3-5. VHF Control AN/ARC-186(V) (Sheet 2 of 2)

CONTROL

FUNCTION

Bandwidth/
MEM
LOAD
(On helicopters
with
panelmounted transceiver). On helicopters with halfsize
remote
control panel, the
memory switch is
labeled LOAD.
Bandwidth
switch is inaccessible.

Three-position switch NB (NARROW) position enables narrowband selectivity WB (WIDE) enables wideband selectivity in the
FM band, momentary MEM
LOAD position allows manually
selected frequency to go into selected preset channel memory.

AM SQUELCH
control (On helicopters
with
panel-mounted
transceiver). (Use
of control is a
maintenance
function).

Screwdriver adjustable potentiometer. Squelch overridden at maximum counterclockwise position,
clockwise rotation increases input
signal required to open the squelch.

3-14

CONTROL

FUNCTION

FM SQUELCH
control (On helicopters
with
panel-mounted
transceiver). (Use
of control is a
maintenance
function).

Screwdriver adjustable potentiometer. Squelch overridden at maximum counterclockwise position,
clockwise rotation increases input
signal required to open the squelch.

Band
LOCKOUT switch (On
helicopters with
panel-mounted
transceiver). (Use
of control is a
maintenance
function).

Screwdriver settable three-position
switch. Center position enables
both AM and FM bands, AM position locks out AM band, FM position locks out FM band. (Band
lockout is indicated by a warning
tone.)

3.7.3 Modes of Operation. Depending on the settings
of the operation controls, the radio set can be used for these
modes of operation:

TM 1-1520-237-10

a. Two-way voice, normal (TR).

b. DF mode check.

b. When voice security system is installed, refer to paragraph 3.11.

(1) Select frequency of station to be used for
homing.

c. Constant monitoring of guard channel 121.5 MHz
only.

(2) Mode select switch - DF.

d. Guard receive and transmit only (EMER).
3.7.3.1 Starting Procedure. Before starting radio set,
check settings of controls that pertain to communication
equipment. With dc power applied, radio set is turned on
with mode selector (Figure 3-5) in any position other than
OFF or EMER.
3.7.3.2 Operational Check. Select mode and communicate with or direction to the ground station on selected
frequencies in low, middle, and high range of applicable
frequency band. Check the action of the volume control and
note that the selected frequencies are heard loud and clear.
Check that adequate sidetone is audible during all transmissions.
a. Communications mode check:
(1) Mode select switch - TR.
(2) Select out-of-band frequency to check
warning. (On helicopters with panelmounted transceivers.)

(3) Frequency control select switch - MAN or
PRE as desired.
(4) Check for homing indication.
c. Squelch disable/tone check.
(1) Select SQ DIS - Check for noise.
(2) Select momentary TONE, check for tone
of about 1000 Hz.
d. Preset channel load.
(1) Mode select switch - TR.
(2) Frequency control select switch - MAN.
(3) Set MHz frequency for desired channel and
rotate PRESET channel to number to be
used with that frequency using channel selector thumbwheel.
(4) LOAD button - Press and release.

(3) Select frequency of station to be used for
check, MAN or PRE as desired.
(4) Communicate with check station.
NOTE
Transmitting with the AN/ARC-186 FM#2
radio may cause the LF-ADF (AN/ARN-89)
bearing pointer to deflect and lose audio
when tuned to a station 400 kHz or below.
Releasing the transmitter key allows the LFADF receiver to return to the normal audio
and bearing indication.
Transmitting with the AN/ARC-186
VHF/AM radio on frequencies from 120
MHz and above may cause the LF-ADF
(AN/ARN-89) bearing pointer to deflect and
lose audio when tuned to a station below
1500 kHz. Releasing the transmitter key allows the LF-ADF receiver to return to the
normal audio and bearing indication.

(5) Repeat steps 3. and 4. for other preset
channels.
3.7.3.3 Operation.
3.7.3.4 TR Mode AM or FM as Applicable.
1. Set OFF-TR-DF switch to TR.
2. Set frequency control select switch to MAN or
PRE.
3. Rotate four MHz selectors to desired frequency
or set PRESET channel number as desired.
3.7.3.5 Emergency Mode AM or FM as Applicable.
1. Mode select switch - TR or DF.
2. Frequency mode selector switch - EMER AM
or FM as applicable.

Change 5

3-15

TM 1-1520-237-10

3.7.3.6 DF (Homing) Mode.
1. Mode select switch - DF.
2. Frequency control select switch - MAN or
PRE.

F
M

LOW
TEST

NORM
HIGH

FAULT
VPA RF IN

LAMP

3.7.3.7 Retransmission Mode. Do a retransmission
check as follows:

OFF

NOTE
Do not disable squelch when retransmit
switches are in retransmit position. Squelch
level is used to key transmitter for retransmission.

AA9242
SA

Figure 3-6. IFM Amplifier Control
1. Establish two base stations at unrelated frequencies.
CONTROL
2. Set appropriate receiver-transmitter to desired
retransmit frequency.

NORM

(Normal power) - 10 watts.

3. Place RADIO TRANSMISSION selector
switch to radios to be used.

HIGH

(High power) - 40 watts.

4. Establish communication between base stations
through aircraft radios.

3.8 RADIO SET AN/ARC-201 (VHF-FM) (IF INSTALLED).

5. Note that selected frequencies are heard loud
and clear and that received audio is present and
clear at each crew station.

Radio set AN/ARC-201 (Figure 3-7) is an airborne, very
high frequency (VHF), frequency modulated (FM), radio
receiving-transmitting set compatible with the Single Channel Ground Airborne Radio Sets (SINGCARS) Electronic
Countermeasures (ECCM) mode of operation. The set provides communications of voice and data, secure or plain
text, and homing over the frequency range of 30 to 87.975
MHz channelized in 25 Khz steps. A frequency offset tuning capability of -10 Khz, -5 Khz, +5 Khz and +10 Khz is
provided in both transmit and receive mode; this capability
is not used in ECCM mode. The set when used in conjunction with the TSEC/KY-58 equipment, is used for receiving
and transmitting clear-voice or X-mode communications.
An additional capability for retransmission of clear-voice
communications allows use of the set as a relay link. During retransmission, when one radio receives a signal, it
sends a keying signal to the second radio and the first radio’s received audio modulates the second radio’s transmitter. Use of the homing capability of the No. 1 FM radio set
provides a steering output to the VSI course deviation
pointer for steering indications. No. 1 VHF-FM receives
power from the dc essential bus through a circuit breaker

3.7.4 Stopping Procedure. Mode Selector - OFF.
3.7.5 IFM Amplifier Control.
NOTE
IFM amplifier control installed in EH-60A
helicopters with ARC-186 as VHF-FM No.
1 (Figure 3-6).

CONTROL

FUNCTION

OFF

(Bypass amplifier) - 10 watts.

LAMP

Tests indicator lights.

TEST

Checks IFM amplifier.

LOW

(Low power) - 2.5 watts.

3-16

FUNCTION

TM 1-1520-237-10

marked NO. 1 VHF-FM. No. 2 VHF-FM receives power
from the No. 1 dc primary bus through a circuit breaker
marked NO. 2 VHF-FM.

CONTROL

FUNCTION

MODE
3.8.1 Antennas.
a. UH The No. 1 VHF-FM communications antenna is
on top of the tail rotor pylon (Figure 3-1). The No. 2
VHF-FM antenna is within the leading edge fairing of the
tail pylon drive shaft cover. The FM homing antennas, one
on each side of the fuselage, are used with FM No. 1 radio
set. The troop commander’s antenna is on the upper trailing
edge of the tail pylon. Refer to Chapter 4, Section 1 for use
of troop commander’s antenna.
b. EH The No. 1 VHF-FM communications antenna is
within the leading edge fairing of the tail pylon drive shaft
cover (Figure 3-1). The No. 2 VHF-FM antenna is under
the nose section. The FM homing antennas, one on each
side of the fuselage, are used with FM No. 1 radio set. the
troop commander’s antenna in on the upper trailing edge of
the tail pylon. Refer to Chapter 4, Section 1 for use of troop
commander’s antenna.
3.8.2 Controls and Functions. Controls for the ARC201 are on the front panel (Figure 3-7). The function of
each control is as follows:
CONTROL

Homing
antennas
selected;
communication
antenna
disconnected. Provides pilot with
steering, station approach and
signal strength indicators.

Single
Channel.
Operating
frequency is selected by PRESET
switch or keyboard entry.

Frequency Hopping. PRESET
switch
positions
1-6
select
frequency parameters.

FH-M

Frequency hopping-master selects
control station as the time standard
for communicating equipment.

PRESET
MAN

Used in single channel mode to
select any operating frequency in
25 Khz increments.

Positions 1-6

In SC mode, preset frequencies are
selected or loaded. In FH or FH-M
mode, frequency hopping nets are
selected.

CUE

Used by a non-ECCM radio to
signal ECCM radio.

FUNCTION

OFF

Primary power off;
Memory battery power ON.

TEST

RT and ECCM modules are tested,
Results: GOOD or FAIL.

SQ ON

RT on with squelch.

SQ OFF

RT on with squelch disabled.

RXMT

RT is receiving. Used as a radio
relay link.

Keyboard loading
frequencies.

preset

LD-V

TRANSEC variable loading is
enabled.

Z-A

Pull and turn switch. (Not an
operational position). Used to clear
the TRANSEC variable.

STOW

HOM

Pull and turn switch. All power
removed. Used during extended
storage.

IFM RF PWR (VHF-FM No. 1 only)
NOTE

This switch is inactive for VHFFM No. 2, leave switch in OFF
position.
OFF

(Bypass amplifier) - 10 watts.

(Low power) - 2.5 watts.

NORM

(Normal power) - 10 watts.

(High power) - 40 watts.

VOL control

Adjust
receiver
comfortable level.

volume

KEYBOARD
Switches 1-9

To key in any frequency, load time
information or offsets.

3-17

TM 1-1520-237-10

IFM RF PWR
NORM HI
LO
OFF

PRESET
2 3 4
5
1
MAN

6
CUE

L E
3

FREQ

ERF
OFST

TIME

CLR

H−Ld
0

MODE
FUNCTION
RXMT
LD
SQ OFF
SQ ON
LD−V
Z−A
TEST
STOW

OFF

Sto
ENT

HOM

SC
FH
FH−M

VOL
AA9243
SA

Figure 3-7. FM Control AN/ARC-201

FUNCTION

b. Two-way voice, secure voice, when TSEC/
KY-58 is installed. Refer to paragraph 3.11.

FREQ

Display current operating frequency during single channel
(manual or preset) operation.

c. Two-way voice, frequency hopping (FH or
FH-M). Secure voice can be used at the same
time if desired.

ERF/OFST

Modify single channel operating
frequency, manually selected or
preset, to include offsets of 65 Khz
or 610 Khz.

CONTROL

TIME

Used to display or change the time
setting maintained within each RT.

STO

Store or enter any frequency into
RT; store a received HOPSET or
LOCKOUT set held in holding
memory.

d. Homing (HOM).
e. Retransmission (Function - RXMT).
3.8.4 Starting Procedure. The radio is capable of operating in any of the modes indicated by the MODE selector switch (Figure 3-7) and retransmission on the FUNCTION switch.
3.8.4.1 Single Channel (SC) Mode.

0/LOAD

Enter zeros; initiate transfer of
ECCM parameters.

1. FUNCTION - SQ ON or SQ OFF.

CLR

Zeroize the display; clear erroneous
entries.

2. MODE - SC.

3.8.3 Mode of Operation. Depending on the setting of
the operating controls, the radio set can be used for these
modes of operations:
a. Two-way voice, normal (SC).

3-18

3. PRESET - MAN.
4. Push FREQ then CLR button. The display will
show all bottom dashes.
5. Enter frequency - 5 digits.

TM 1-1520-237-10

6. Push STO button. The display will flash once
to acknowledge correctly entered frequency.
7. ICS transmitter selector - Position 1 (FM No.
1), or position 4 (FM No. 2).
8. Radio push-to-talk switch - Press to talk; release to listen.
3.8.4.2 Enter Frequency into PRESET.
1. FUNCTION - LD.

deviation from the course to the transmitting station.
c. Station passage will be indicated by course
deviation change and CIS MODE SEL
NAV switch light going out and HDG
switch light going on.
3.8.5 Retransmission (RXMT). Retransmission permits helicopter to be used as an airborne relay (Figure
3-12).

2. PRESET - Desired number 1 to 6.

1. FUNCTION - RXMT.

3. MODE - SC.

2. Frequency(s) - Select.

4. Push FREQ then CLR button. The display will
show all bottom dashes.

3. RADIO RETRANSMISSION selector switch
- Set to radios used.

5. Enter frequency - 5 digits.

4. Establish communications between each relay
in helicopter and its counterpart radio at the
terminal station by using appropriate ICS
TRANS selector. If audio monitoring is desired,
adjust VOL control for suitable output.

6. Push STO. The display will flash once.
7. Repeat steps 1. through 6. for each desired preset channel.

3.8.6 Stopping Procedure. FUNCTION - OFF.

3.8.4.3 Frequency Hopping (FH or FN-M) Mode.
3.9 DELETED.
1. MODE - FH or FH-M.
2. PRESET - Select net (1-6).
3. FUNCTION - SQ ON or SQ OFF.
3.8.4.4 Homing (HOM) Mode (FM No. 1 only).
1. Enter or select frequency - MAN or PRESET.
2. MODE - HOM.
3. MODE SEL - FM HOME.
4. CIS MODE SEL - NAV.
5. Observe homing indicators on vertical situation
indicator (VSI) (Figure 3-30). These are:
a. FM navigation (NAV) flag will move from
view, and will come into view if the received signal is too weak.
b. A steering (course indicator) pointer moves
either left or right about 5° to indicate any

3.10 RECEIVER-TRANSMITTER RADIO, RT-1518C/
ARC-164(V).
Receiver-Transmitter Radio, RT-1518C/ARC-164(V)
(Figure 3-8) is an airborne, ultra-high frequency (UHF),
amplitude-modulated (AM), radio transmitting-receiving
(transceiver) set with Have Quick II (provides Have Quick
I and Have Quick II capabilities). Have Quick provides
electronic counter countermeasures (ECCM) which use a
frequency hopping scheme to change channels multiple
times per second. The RT-1518C/ARC-164(V) contains a
multichannel, electronically tunable main transmitter and
receiver, and a fixed-tuned guard receiver. The main transceiver operates on any one of 7,000 channels, spaced in
0.025 MHz units in the 225.000 to 399.975 MHz UHF
military band. The guard receiver is always tuned to
243.000 MHz. The radio is primarily used for voice communications. An additional capability for retransmission allows use of the radio as a relay link. The radio interfaces
with the Doppler/GPS Navigation Set (DGNS) AN/ASN128B. Power to operate the ARC-164(V) radio is from the
dc essential bus through a circuit breaker marked UHF
AM.

Change 10

3-19

TM 1-1520-237-10

FREQUENCY /
STATUS
INDICATOR

CHANNEL
INDICATOR

FREQUENCY
COVER
CH

TEST
DISPLAY
BUTTON

FREQ

CHAN

CHANNEL
SELECTOR

TEST
DISPLAY

FREQUENCY
SELECTOR 2

U
H
F

FREQUENCY
SELECTOR 1

STATUS
BUTTON

STATUS

3
2

VOL
MAIN
OFF

PRESET
MNL
GRD

BOTH
ADF

SQUELCH
T

TONE

OFF

FREQUENCY
SELECTOR 5
FREQUENCY
SELECTOR 4

MNL−PRESET−GRD
SELECTOR
MODE
SELECTOR

FREQUENCY
SELECTOR 3

VOLUME
CONTROL

SQUELCH
SWITCH

A
ZERO
SWITCH

MAIN
SQUELCH
CONTROL

FILL
CONNECTOR

ZERO
F
I
L
L

MN SQ
LOAD

LOAD
BUTTON

GD SQ

GUARD
SQUELCH
CONTROL

AB2433
SA

Figure 3-8. UHF Control, AN/ARC-164(V)

3-20

Change 10

TM 1-1520-237-10

3.10.1 Antenna.

CONTROL

FUNCTION

WARNING

TEST DISPLAY
button

Lights all segments of the
frequency/status and CHAN indicator. Also used with T TONE
switch for manual clock start.

The antenna radiates electromagnetic
waves at dangerous frequencies. Comply
with the requirements of AR 40-583 before using this equipment.

Frequency/status
indicator

Displays the individual frequency
selector settings or any of the following operator prompts:

The UHF-AM antenna is under the fuselage transition
section (Figure 3-1) or EH under the fuselage below the
copilot’s seat. The antenna provides a path for both the
transmitted and received UHF communication signals. The
EH-60 AN/ALQ-151(V) mission package has two
UHF-AM radios that utilize the existing fuselage transition
section conformal antenna for the voice link. The mission
package also includes the data link antenna under the fuselage where the cargo hook would normally be installed.
3.10.2 Tunable Diplexer. EH The tunable diplexer
(TD-1336/A) is connected between the antenna and the output of the ARC-164(V). When properly tuned, the diplexer
acts as a bridge network isolating signals of similar frequency which share the same antenna. The diplexer allows
the guard channel in the ARC-164(V) guard receiver to be
monitored while other frequencies in the main transmitterreceiver are being used.
3.10.3 Controls and Functions. Controls for the
ARC-164(V) are on the front panel of the unit (Figure 3-8).
The function of each control is as follows:

CONTROL
CHAN indicator

Preset channel
selector

FUNCTION
Displays selected channel when
MNL-PRESET-GRD selector is
set to PRESET, or displays selected memory location when radio
is in the MWOD load mode or
FMT change mode.
Selects one of 20 preset channels.
Also selects the desired memory location when radio is in the MWOD
load mode (20-14), manual TOD
entry mode (1), or FMT change
mode (20-5).

REMOTE

Not used.

VER/OP

Indicates radio is in normal operating mode.

M-LOAD

Indicates radio is in MWOD load
mode.

ERASE

Indicates radio is in MWOD erase
mode.

FMTCNG

Indicates radio is in Frequency
Management Training Change
mode.

FILL

Indicates a keyfill device is connected to the front panel FILL connector.

WOD OK

Indicates a valid WOD was successfully received from the keyfill
device.

BAD

Indicates no WOD or a bad parity
WOD was received from the keyfill
device.

STATUS button

Initiates an alternate display on the
frequency/status and CHAN indicator for five seconds.

Frequency
selector 1 (A, 3, 2
HQ only)

Selects 100s digit of frequency (either 2 or 3) in MHz in the single
frequency mode. Selects the desired
WOD elements or net number in AJ
(Have Quick) mode.

Selects AJ mode (Figure 3-8).

Allows manual selection of frequencies in the 300 MHz range
(3XX. XX).

Change 10

3-21

TM 1-1520-237-10

CONTROL
2

FUNCTION
Allows manual selection of frequencies in the 200 MHz range
(2XX. XX).

Frequency
selector 2

Selects 10s digit of frequency (0
through 9) in MHz. Selects the desired WOD elements or net number
in AJ (Have Quick) mode.

Frequency
selector 3

Selects units digit of frequency (0
through 9) in MHz. Selects the desired WOD elements or net number
in AJ (Have Quick) mode.

Frequency
selector 4

Selects tenths digit of frequency (0
through 9) in MHz. Selects the desired WOD elements or net number
in AJ (Have Quick) mode.

Frequency
selector 5

Selects hundredths and thousandths
digits of frequency (.00, .25, .50, or
.75) in MHz. Selects the desired
WOD elements or net number in AJ
(Have Quick) mode.

Mode selector

FUNCTION

T TONE switch

Three position toggle switch:
middle position is off, and T and
TONE are spring loaded. When
placed in the TONE position, transmits a 1,020 Hz DF tone on the selected frequency and stops when
switch is released. If radio TOD
clock is started, the TONE position
transmits the TOD message data
followed by the DF tone. When
placed in the T position, enables reception of TOD message for one
minute. The T TONE switch also
initiates manual TOD clock start,
loads and erases MWOD elements,
and loads FMT-net operating frequencies when radio is in the respective MWOD operating mode.

VOL control

Adjusts volume.

SQUELCH
switch

Disables and enables squelch of
main receiver.

MNL-PRESETGRD selector

Selects method of frequency selection:

Selects operating mode function:

OFF

Turns power off.

MAIN

Enables main receiver and transmitter.

BOTH

Enables main receiver, transmitter,
and guard receiver.

ADF

Not used.

3-22

CONTROL

Change 10

MNL

Allows manual selection of frequency using the five frequency selectors.

PRESET

Allows selection of frequency from
preset channels (1-20) using the
channel selector. Along with
LOAD switch, also used when programming the 20 preset channels.

GRD

Automatically tunes the radio main
receiver and transmitter to the
guard frequency (243.000 MHz),
and disables the guard receiver.

MN SQ control

Adjusts level of squelch for main
receiver.

ZERO switch

Erases all MWOD elements.

LOAD
switch

Loads frequency data displayed on
the frequency/status indicator into
preset channels 1-19 as selected by
channel selector. Preset channel 20
is reserved for loading MWOD operating mode data (220.0XX).

TM 1-1520-237-10

CONTROL
GD SQ control
Fill connector

FUNCTION
Adjusts level of squelch for guard
receiver.
Connects a keyfill device to radio
for automatic loading of MWOD.

3.10.4 Modes of Operation. The radio has three different methods of frequency selection as determined by the
position of the MNL-PRESET-GRD selector. An explanation of these three positions is given in paragraph 3.10.3. A
procedure for programming the 20 preset channel numbers
to the desired frequencies is given in paragraph 3.10.5.7.
The radio can operate in MAIN, BOTH, or ADF.
3.10.4.1 Normal Mode. The normal mode allows twoway voice communications.
3.10.4.2 Secure Speech (x-mode) Mode. The secure
speech mode allows secure two-way voice communications. Operation is identical to the normal mode.
3.10.4.3 ADF Mode. This mode is not normally used.
Transmission from the radio is normally not possible with
the mode selector in the ADF position.
3.10.4.4 1,020 Hz Tone Signal Mode. This mode allows transmission of a 1,020 Hz tone.
3.10.4.5 Guard Channel Mode. The BOTH position
allows use of the main receiver while monitoring the guard
receiver. The guard receiver is always tuned to 243.000
MHz.
3.10.4.6 Retransmit Mode. Retransmission permits the
helicopter to be used as an airborne relay link. Two RTs are
required (Figure 3-12) for operation as a relay unit. Operation in the retransmit mode is identical to the normal mode.
3.10.4.7 Have Quick/Anti-Jam Mode. The Have
Quick II system provides jam resistant (anti-jamming) capability through frequency hopping. Frequency hopping is
when the frequency being used for a given channel is automatically changed at a rate common to the transmitter
and receiver. The jam resistance of the system is due to the
automatic frequency changing and the pseudorandom pattern of frequencies used. Certain criteria are necessary for
successful system operations. These are:
1.

Common frequencies.

Time synchronization.

Common hopping pattern and rate.

Common net number.

The common frequencies are programmed into all Have
Quick radios. Time synchronization is provided via UHF
radio and/or hardware by an external time distribution system. A time-of-day (TOD) signal must be received from the
time distribution system each time the radio is turned on.
The hopping pattern and hopping rate are determined by
the operator inserted word-of-day (WOD). In the AJ mode,
a communications channel is defined by a net number. In
addition to these Have Quick I capabilities, Have Quick II
provides two new frequency tables; multiple word of day
(MWOD); MWOD erase capability; Frequency Management Training net (FMT net) in addition to existing Training net (T net); and operational date information as part of
TOD for proper WOD initialization.
3.10.4.7.1 Word Of Day (WOD). WOD is entered by
using one or more of the six preset channels (15-20).
3.10.4.7.2 Multiple Word Of Day (MWOD). MWOD
allows loading of up to six WODs either manually or automatically (maintenance can load the radio using a keyfill
device). Each WOD contains a unique date code that is
entered in memory location (channel) 14. The four MWOD
operating modes: VER/OP, M-LOAD, ERASE, and
FMTCNG are used to initiate various programming functions.
3.10.4.7.3 Time Of Day (TOD).
NOTE
Automatic TOD is provided from the
Doppler/GPS Navigation Set (DGNS) AN/
ASN-128B.
TOD allows radios to operate together in the AJ mode.
Transmission and reception are possible in both single frequency and AJ modes. Slightly garbled but otherwise acceptable communications indicate the radios have drifted
out of synchronization. A time update/resynchronization
corrects this. If single WOD is used, an operational date is
not necessary.
3.10.4.7.4 TOD Clock Manual Start. When TOD from
a UTC (Universal Coordinated Time) source is not available, the TOD clock can be manually started and used in
the AJ mode. The manually started TOD clock is set to a
time completely independent of UTC. Other radios may
also communicate using this uncoordinated time once the
time has been transmitted to all radios in the net.

Change 10

3-23

TM 1-1520-237-10

3.10.4.7.5 Net Number. The net number programs the
entry point in the AJ frequency hopping pattern, allowing
multiple radio net operations using a common WOD and
TOD without interfering with each other. Selecting A with
frequency selector 1 places the radio in AJ mode and programs the radio to use the net number selected by frequency
selectors 2, 3, and 4. The net number begins with A and is
followed by three digits (000 to 999).
Operational Net Numbers. The last two digits designate the frequency table being used. Net numbers
ending in 00 select the original A-net and B-net frequency table. Net numbers ending in 25 select the
new NATO/Europe frequency table. Net numbers
ending in 50 select the new non-NATO/Europe frequency table. Net numbers ending in 75 are reserved
for future use and will generate an invalid net alarm
(pulsating warning tone).
Training/FMT Net Numbers. T-net and FMT-net
training numbers are available for training purposes.
In the T-net training mode, the radio hops among five
frequencies loaded in with the WOD. In the FMT-net
training mode, the radio hops among sixteen frequencies.
3.10.4.7.6 Conference Capability. In the AJ mode,
the radio is able to receive and process two simultaneous
transmissions on the same net. In a conference net, the
second transmitting radio automatically shifts its transmission frequency 25 kHz when it monitors a transmission on
the primary net frequency. The wide band receiver reads
both transmissions without the interference normally associated with two radios transmitting simultaneously on the
same frequency. Conference capability is enabled or disabled by the last two digits of the WOD element loaded in
memory location (channel) 19. If the WOD element ends
with 00 or 50, conferencing is enabled. If the WOD element ends with 25 or 75, conferencing is disabled. When
operating in secure speech mode, conferencing is automatically disabled. If operating in AJ mode on a single element
WOD (memory location (channel) 20 only), conferencing
is enabled by default. For training mode operation, conferencing is always enabled.
3.10.5 Operation.
3.10.5.1 Normal Mode.
1. Mode selector - MAIN.
2. MNL-PRESET-GRD selector - As desired.
3. Frequency selector/channel selector - As desired.

3-24

Change 10

4. ICS transmitter selector switch - Position 2.
5. Establish communication by keying transmitter
and speaking into microphone. Release to listen
and adjust volume for a comfortable level.
3.10.5.2 Secure Speech Mode.
1. Refer to paragraph 3.11 for voice security system procedures.
2. Refer to 3.10.5.1 for normal mode procedures.
3.10.5.3 1,020 Hz Tone Signal Mode.
1. Mode selector - MAIN.
2. MNL-PRESET-GRD selector - As desired.
3. Frequency selector/channel selector - As desired.
4. T TONE switch - Press to TONE to transmit
the 1,020 Hz signal.
3.10.5.4 Guard (Emergency) Channel Mode.
Guard frequency only:
1. Mode selector - BOTH.
2. MNL-PRESET-GRD selector - GRD (main
receiver and transmitter are both tuned to
243.000 MHz).
Guard frequency and main (receiver/transmitter) frequency:
1. Mode selector - BOTH.
2. MNL-PRESET-GRD selector - MNL or
PRESET.
3. Frequency selector/channel selector - As desired for main receiver/transmitter.
4. ICS transmitter selector switch - Position 2.
5. Establish communication by keying transmitter
and speaking into microphone. Release to listen
and adjust volume for a comfortable level.

TM 1-1520-237-10

3.10.5.5 Retransmit Mode.
1. RADIO TRANSMISSION selector - Set to
desired radios.
2. Mode selector - MAIN.
3. MNL-PRESET-GRD selector - As desired.
4. Frequency selector/channel selector - As desired.
5. ICS transmitter selector switch - Position 2.
6. Establish communication between each relay
radio in helicopter and its counterpart radio link
terminal station.

3.10.5.6.1 Manual WOD Loading.
NOTE
If power is lost, or if channel 20 is selected
when the MNL-PRESET-GRD selector is
in the PRESET position, reinitialization of
all WOD elements is necessary.
Use of all channels may not be necessary.
Any unused channels can be used to store
selected preset frequencies.
1. Mode selector - MAIN.
2. MNL-PRESET-GRD selector - PRESET.
3. Channel selector - Channel 20.

3.10.5.6 Have Quick/AJ Mode.
1. Mode selector - MAIN or BOTH.
2. WOD or MWOD - Load per paragraph
3.10.5.6.1 or 3.10.5.6.2.
3. TOD - Synchronize per paragraph 3.10.5.6.4.
4. Verify/operate mode (MWOD only) - Select
per paragraph 3.10.5.6.3.
5. MNL-PRESET-GRD selector - MNL.
NOTE
A steady warning tone is heard when the AJ
mode is selected and a TOD or a valid WOD
has not been entered.
The A cannot be stored in preset channel
memory. If loading a net number into a preset channel is attempted, the A is accepted
as a 3.

4. Frequency selectors - Set desired WOD element.
5. LOAD button (under frequency cover) - Press.
6. If needed, repeat steps 3 through 5 for channels
19 - 15.
7. Initialization: a single beep is heard on each
channel with a WOD element with the exception of the final WOD element which has a
double beep. This indicates the end of WOD
elements and initialization is complete.
8. MNL-PRESET-GRD selector - MNL.
3.10.5.6.2 Manual MWOD Loading.
NOTE
If the frequency selectors are not used within
five seconds, the frequency/status indicator
reverts to the M-LOAD display. Pressing
the STATUS button allows reviewing of the
frequency settings.

6. Frequency selector 1 - A.
7. Frequency selectors 2, 3, 4 - Enter net number.
A pulsating warning tone is heard if an invalid
operating net is selected.
8. Establish communication by keying transmitter
and speaking into microphone. Release to listen
and adjust volume for a comfortable level.

When the current operational date is updated
in the radio at midnight (Greenwich Mean
Time), the radio automatically reinitializes
with the WOD having the same date code.
1. Mode selector - MAIN.
2. MNL-PRESET-GRD selector - PRESET.

Change 10

3-25

TM 1-1520-237-10

3. Channel selector - Channel 20.
4. Frequency selectors - Set 220.025 MHz to select MWOD load mode.

4. Channel selector - Memory location (channel)
20, then 19, and then 20. Single beep verifies
an MWOD with a matching date code is stored
in memory. If a single beep is not heard, the
selected date code is not stored in memory.

5. LOAD button (under frequency cover) - Press.
5. Proceed to paragraph 3.10.5.6.4 for TOD.
6. MNL-PRESET-GRD selector - MNL.
7. Frequency selectors - Set WOD element.
8. T TONE switch - Press to TONE, then release
and note a single beep to ensure WOD element
is entered.
9. Channel selector - Select next lower memory
location.
10. Repeat steps 7 through 9 for loading remaining
WOD elements in memory locations (channels)
19 - 15.
11. Channel selector - Channel 14 (date tag).
12. Frequency selectors - Set to applicable date
code: XAB. XXX (AB represents the day-ofmonth (01 to 31); Xs do not need a data entry).
If two or more WODs have the same date code,
the radio recognizes the last one entered.
13. T TONE switch - Press to TONE, then release
(note a double beep). One complete WOD with
date code has been successfully entered.
14. Reselect channel 20 and repeat steps 7 through
13 to load more WODs. If power is removed
from the radio, MWOD data is not lost. All
data remains in memory unless erased. Refer to
the operational date load or change procedures
in paragraph 3.10.5.6.4 for MWOD data recall.
MWOD Date Code Verify:
1. Select MWOD verify/operate mode per paragraph 3.10.5.6.3.
2. MNL-PRESET-GRD selector - MNL.
3. Frequency selectors - Select date code to be
verified: XAB. XXX (AB represents the dayof-month (01 to 31); Xs do not need a data
entry).

3-26

Change 10

3.10.5.6.3 MWOD Operating Mode Selection.
1. Channel selector - Channel 20.
2. MNL-PRESET-GRD selector - PRESET.
3. Frequency selectors - As desired (220.000 MHz
for verify/operate mode; 220.025 MHz for
MWOD load mode; 220.050 MHz for MWOD
erase mode; or 220.075 MHz for frequency
management training change mode).
4. LOAD button (under frequency cover) - Press
to enable selected mode. The appropriate operating mode (M-LOAD, ERASE, or
FMTCNG) is displayed on the frequency/
status indicator (the STATUS button must be
pressed to display the VER/OP mode).
5. MNL-PRESET-GRD selector - MNL.
MWOD erase:
1. Perform steps 1 through 5 to select MWOD
erase mode.
2. T TONE switch - Press to TONE. All
MWODs are now cleared (WOD data and net
number information is not cleared).
3. MNL-PRESET-GRD selector - PRESET.
4. Frequency selectors - Set 220.000 MHz for
verify/operate mode.
5. LOAD button (under frequency cover) - Press
to enable VER/OP mode (the STATUS button
must be pressed to display the VER/OP on the
frequency/status indicator for five seconds).
6. MNL-PRESET-GRD selector - MNL.

TM 1-1520-237-10

Alternate method for MWOD erase:
1. ZERO switch (under frequency cover) - Press,
then release. ERASE is displayed (all MWODs
are now erased). The ZERO switch only erases
the MWOD data. WOD data and net number
information are not erased.

1. MNL-PRESET-GRD selector - As desired.
2. Frequency selectors/channel selector - Select
predesignated frequency for TOD transmission.
NOTE
The first TOD received within one minute of
the TOD request is accepted.

FMT change:
1. Perform steps 1 through 5 to select MWOD
FMT change mode.
2. Channel selector - Set desired memory location
(channels 20-5).
NOTE
If the frequency selectors are not used within
five seconds, the frequency/status indicator
reverts to the current operating mode display
(FMTCNG).
3. Frequency selectors - Select frequency change.
4. T TONE switch - Press to TONE, then release
(note an audible tone).
5. Repeat steps 2 through 4 until all frequency
changes are loaded.
6. Channel selector - Channel 20.
7. MNL-PRESET-GRD selector - PRESET.
8. Frequency selectors - Set 220.000 MHz for
verify/operate mode.
9. LOAD button (under frequency cover) - Press
to enable VER/OP mode (the STATUS button
must be pressed to display the VER/OP on the
frequency/status indicator for five seconds).
3.10.5.6.4 Time of Day (TOD).
Request/Receive TOD:
NOTE
When the radio is turned on, the first TOD
received is accepted. Subsequent messages
are ignored.

3. T TONE switch - Press to T then release.
4. If time is not being automatically beaconed, request TOD from another station on the operating network. Beacons typically transmit TOD
every ten seconds.
5. TOD may be received on the main or guard
receiver in single frequency or AJ mode. Two
momentary tones (1,667 Hz-high tone and
1,020 Hz-low tone) are heard when the TOD
synchronization signal is received.
Transmit TOD:
1. MNL-PRESET-GRD selector - As desired.
2. Frequency selector/channel selector - As desired.
NOTE
The radio will not transmit while the T
TONE switch is in the T position.
3. T TONE switch - Press to TONE then release.
Two momentary tones (1,667 Hz-high tone and
1,020 Hz-low tone) are heard when the TOD
synchronization signal is transmitted.
TOD Resynchronization:
1. T TONE switch - Press to T then release.
2. Frequency selectors - Select any normal frequency and request a TOD. The first TOD signal is heard within one minute of selecting T is
accepted. A momentary 1,667 Hz tone is heard
when the TOD signal is received. TOD resynchronization should be performed using the
single frequency mode.

Change 10

3-26.1

TM 1-1520-237-10

TOD Clock Manual Start:

3.10.5.7 Preset Channel Loading.
NOTE

NOTE
The new TOD is arbitrary and is not synchronized to UTC or to any other radio.
A manual TOD start clears out a previously
loaded TOD.
1. T TONE switch - Press to T and hold, while
simultaneously pressing the TEST DISPLAY
button.
2. T TONE switch - Release prior to releasing the
TEST DISPLAY button to prevent inadvertent
Have Quick time loading.
Operational Date Load or Change:

Channels 14 through 20 can be reserved for
loading single WOD and MWOD data. Any
channel not used for WOD or MWOD can
be used as a preset channel.
The A cannot be stored in preset channel
memory. If loading of a net number (AXX.
XXX) into a preset channel is attempted, the
A is accepted as a 3.
Set 20 preset channel numbers to desired frequencies as
follows (Figure 3-8):
1. Mode selector - MAIN or BOTH.
2. MNL-PRESET-GRD selector - PRESET.

1. Mode selector - MAIN.

3. Frequency selectors - As desired.

2. MNL-PRESET-GRD selector - PRESET.

4. Channel selector - As desired.

3. Channel selector - Channel 20.

5. LOAD button - Press, then release.

4. Frequency selectors - Set 220.025 MHz to select MWOD load mode.

6. Using a pencil, record frequency selected for
channel on the card located on the front panel.

5. LOAD button (under frequency cover) - Press.

7. Repeat steps 3 through 6 to load additional preset channels.

6. MNL-PRESET-GRD selector - MNL.
3.10.6 Stopping Procedure. Mode selector - OFF.
7. Channel selector - Channel 1.
3.11 VOICE SECURITY SYSTEM.
8. Frequency selectors - Set to applicable date
code: XAB. XXX (AB represents the day-ofmonth (01 to 31); Xs do not need a data entry).
9. T TONE switch - Press to TONE, then release
and note a tone (the date code has been successfully entered).
10. Frequency selectors - Set to 220.000 MHz to
select verify/operate mode.
11. LOAD button (under frequency cover) - Press,
then release. Radio is now in verify/operate
mode.
12. STATUS button - Press, VER/OP is displayed
on the frequency/status indicator for five seconds.
3-26.2

Change 10

3.11.1 Deleted. Figure 3-9 deleted.
3.11.2 TSEC/KY-58.
a. Description. A complete description of the TSEC/
KY-58 can be found in TM 11-5810-262-10. This voice
security equipment is used with the FM1, FM2 and
UHF-AM radio to provide secure two-way communication.
The equipment is controlled by the Remote Control Unit
(RCU) (Z-AHP) mounted in the lower console. The
POWER switch must be in the ON position, regardless of
the mode of operation, whenever the equipment is installed.
b. Controls and functions. Figure 3-10.
c. Operating procedures.

TM 1-1520-237-10

NOTE
To talk in secure voice, the KY-58 must be
9loaded9 with any number of desired variables.
(1) Secure voice procedures.
(a) MODE switch - OP.

(b) FILL switch - Set to the storage register
which contains the crypto-net variable
(CNV) you desire.
(c) POWER switch - ON.
(d) PLAIN, C/RAD switch - C/RAD1.
(e) DELAY switch - Down unless the signal
is to be retransmitted.

Change 10

3-26.3/(3-26.4 Blank)

TM 1-1520-237-10

NOTE
At this time a crypto alarm, and background
noise, in the aircraft audio intercom system
should be heard.

(b) Several beeps should now be heard in your
headset. This means that the 9old9 CNV is
being replaced by a 9new9 CNV.
(c) Using this 9new9 CNV, the net controller
will ask you for a 9radio check.9

To clear alarm:
(f) PTT (push-to-transmit) switch - Press and
release.
NOTE
When operating in either secure or clear
(plain) voice operations the volume must be
adjusted on the aircraft radio and intercom
equipment to a comfortable operating level.
(2) Clear voice procedures:
(a) POWER switch - ON.
(b) PLAIN, C/RAD switch - PLAIN.
NOTE
Instructions should originate from the Net
Controller or Commander as to when to zeroize the equipment.
(3) Zeroing procedures.
(a) POWER switch - ON.
(b) Spring-loaded ZEROIZE switch - Activate and release. This will zeroize all positions (1-6). The equipment is now zeroized
and secure voice communications are no
longer possible.
(4) Automatic remote keying procedures.
NOTE
Automatic remote keying (AK) causes an
9old9 crypto-net variable (CNV) to be replaced by a 9new9 CNV. Net controller simply transmits the 9new9 CNV to your KY-58.

(d) After the 9radio check9 is completed, the
net controller instructions will be to resume
normal communications. No action should
be taken until the net controller requests a
9radio check.9
(5) Manual remote keying procedures.
(a) The net controller will make contact on a
secure voice channel with instructions to
stand by for a new crypto-net variable
(CNV) by a manual remote keying (MK)
action. Upon instructions from the net controller:
1 Set the Z-AHP FILL switch to position 6. Notify the net controller by radio, and stand by.
2 When notified by the net controller, set
the Z-AHP MODE switch to RV (receive variable). Notify the net controller, and stand by.
3 When notified by the net controller, set
the Z-AHP FILL switch to any storage position selected to receive the
new CNV (May be unused or may
contain the variable being replaced).
Notify the net controller, and stand by.
NOTE
When performing Step 3, the storage position (1 through 6) selected to receive the new
CNV may be unused, or it may contain the
variable which is being replaced.
(b) Upon instructions from the net controller:
1 Listen for a beep on your headset.
2 Wait two seconds

(a) The net controller will use a secure voice
channel, with directions to stand by for an
AK transmission. Calls should not be made
during this standby action.

3 Set the RCU MODE switch to OP
4 Confirm

3-27

TM 1-1520-237-10

PLAIN
C / RAD

MODE
OP
LD

KY
58
R

RV
DELAY

C
U

E
O
I
Z
E

1
2

3
4

5
6

6
FILL

POWER

REMOTE CONTROL UNIT (RCU) (Z−AHP)
CONTROL

FUNCTION

1. ZEROIZE SWITCH TWO−
POSITION MOMENTARY
TOGGLE UNDER SPRING
LOADED COVER

ZEROIZES THE KY−58; CLEARS ALL ENCODING IN
THE SYSTEM.

2. DELAY SWITCH TWO−
POSITION TOGGLE

UP WHEN SIGNAL IS TO BE RETRANSMITTED.

3. PLAIN−C / RAD SWITCH
ROTARY TWO−POSITION
SELECTOR SWITCH

IN THE PLAIN POSITION, PERMITS NORMAL (UNCI−
PHERED) COMMUNICATIONS ON THE ASSOCIATED
FM RADIO SET.
IN THE C / RAD POSITION, PERMITS CIPHERED COM−
MUNICATIONS ON THE ASSOCIATED RADIO SET.

4. C / RAD2 SWITCH STOP

LOCATION STOP FOR C / RAD2 ON FRONT PANEL.

5. FILL SWITCH SIX POSITION
ROTARY SWITCH

PERMITS PILOT TO SELECT ONE OF 6 STORAGE
REGISTERS FOR FILLING.

6. MODE SWITCH THREE
POSITION ROTARY

IN THE OP POSITION KY−58 NORMAL OPERATING,
IN THE LD POSITION FOR FILLING.
IN THE RV POSITION KY−58 IN RECEIVE−VARIABLE,
FILLED FROM ANOTHER EXTERNAL SOURCE.

7. POWER ON SWITCH TWO
POSITION TOGGLE

CONNECTS POWER TO THE ASSOCIATED TSEC / KY−58
CIPHER EQUIPMENT IN THE ON (FORWARD) POSI−
TION, AND DISCONNECTS POWER FROM THE EQUIP−
MENT IN THE OFF (AFT) POSITION. TURNS ON POWER
TO TSEC / KY−58.
AA0524
SA

Figure 3-10. Voice Security Equipment
(c) If the MK operation was successful, the net
controller will now contact you via the new
CNV.
(d) If the MK operation was not successful,
the net controller will contact you via clear
voice (plain) transmission; with instructions to set your Z-AHP FILL selector
switch to position 6, and stand by while the
MK operation is repeated.
(6) It is important to be familiar with certain KY-58
audio tones. Some tones indicate normal operation, while other indicate equipment malfunction. These tones are:

3-28

(a) Continuous beeping, with background
noise, is cryptoalarm. This occurs when
power is first applied to the KY-58, or
when he KY-58 is zeroized. This beeping
is part of normal KY-58 operation. To clear
this tone, press and release the PTT button
on the Z-AHQ (after the Z-AHQ LOCAL
switch has been pressed). Also the PTT can
be pressed in the cockpit.
(b) Background noise indicates that the KY-58
is working properly. This noise should occur at TURN ON of the KY-58, and also
when the KY-58 is generating a cryptovariable. If the background noise is not heard

TM 1-1520-237-10

GPS

SA
/
AS

GPS

UHF
T
R
A/
NH
SO
EP
CS
E
T

FM2

FM1

RADIO RETRANSMISSION

K
Y
−
5
8

FM 1 / FM 2
FM 1 / VHF

C
O
M
S
E
C

FM 1 / UHF

FM 2 / UHF
FM 2 / VHF
VHF / UHF

OFF

MODE
SELECTOR
AA9244A

AA0359

Figure 3-11. Remote Fill Panel
at TURN ON, the equipment must be
checked out by maintenance personnel.
(c) Continuous tone, could indicate a 9parity
alarm.9 This will occur whenever an empty
storage register is selected while holding
the PTT button in. This tone can mean any
of three conditions:
1 Selection of any empty storage register.
2 A 9bad9 cryptovariable is present.

Figure 3-12. Retransmission Control Panel
filled storage register is selected, this
tone will be heard. Normal use (speaking) of the KY-58 is possible.
2 When the KY-58 has successfully received a cryptovariable, this tone indicates that a 9good9 cryptovariable is
present in the selected register.
3 When you begin to receive a ciphered
message this tone indicates that the
cryptovariable has passed the 9parity9
check, and that it is a good variable.

3 Equipment failure has occurred. To
clear this tone, follow the 9Loading
Procedures9 in TM 11-5810-262-10. If
this tone continues, have the equipment checked out by maintenance personnel.

(f) A single beep, when the RCU is in TD
(Time Delay) occurring after the 9preamble9 is sent, indicates that you may begin speaking.

(d) Continuous tone could also indicate a cryptoalarm. If this tone occurs at any time
other than in (c) above, equipment failure
may have occurred. To clear this tone, repeat the 9Loading Procedures9 in TM 115810-262-10. If this tone continues, have
the equipment checked out by maintenance
personnel.

(g) A single beep, followed by a burst of noise
after which exists a seemingly 9dead9 condition indicates that your receiver is on a
different variable than the distant transmitter. If this tone occurs when in cipher text
mode: Turn RCU FILL switch to the CNV
and contact the transmitter in PLAIN text
and agree to meet on a particular variable.

(e) Single beep, when RCU is not in TD (Time
Delay), can indicate any of the three normal conditions:

3.11.2.1 KY-58 Remote Fill. A remote fill panel (Figure 3-11) allows a single crew member to load COMSEC
variables into each of the three KY-58’s from the pilots
side lower center console, FM-1 and FM-2 AN/ARC-201
TRANSSEC and HOPSET codes can be loaded from the
same panel.

1 Each time the PTT button is pressed
when the KY-58 is in C (cipher) and a

Change 8

3-29

TM 1-1520-237-10

3.12 RADIO RETRANSMISSION CONTROL.
Control of retransmission is through a switch panel (Figure 3-12) on the lower console. The position of the switch
determines which radio set pairs will be used when the
corresponding FM and VHF radio function and VOL
switches (not required for UHF) are placed to RETRAN.
Operation of the retransmission control is included with the
operating procedures of each radio set where applicable.
The retransmission control is only a means of directing the
audio output of a receiver to the audio input of a transmitter
through switching.
3.12A HF RADIO SET AN/ARC-220.

3.12A.1 Antenna. The tubular antenna element extends
from the left side of the transition area to a point just forward of the hinged tailcone section, and is supported by
four masts. RF energy is supplied to the antenna through
the forward mast (Figure 3-1).
3.12A.2 Controls and Functions. The radio is controlled by a control display unit (CDU) located in the lower
console (Figure 3-12.1). The function of each control and
display is as follows:

CONTROL/
DISPLAY
CURSOR keys.

Position the cursor in the direction
of the arrow on the key.

Display screen

Used
to
display
system
information, and enter data or
commands in radio.

Line select keys

Function
display.

Brightness keys

Changes display screen brightness.

Net selector switch

Selects programmed operating net.
The + position allows additional
nets to be selected using the
VALUE keys.

DATA connector

Fills radio with preprogrammed
data, required for all modes except
MAN.

KEY connector

Used to load ALE and ECCM
presets.

WARNING
Make sure that no personnel are within 3
feet of the HF antenna when transmitting,
performing radio checks or when in ALE
mode. Do not touch the RF output terminal on the antenna coupler, the insulated
feed through, or the antenna itself while
the microphone is keyed (after the tuning
cycle is complete) or while the system is in
transmit self-test. Serious RF burns can
result from direct contact with the above
criteria.
a. The AN/ARC-220 HF transceiver provides long
range communications. The HF radio receives and transmits on any one of 280,000 frequencies spaced at 100 Hz
steps on the high frequency (HF) band. The HF radio has a
frequency range of 2.0000 - 29.9999 MHz. Preset nets can
be manually programmed by the pilot, or loaded with a data
transfer device (DTD). Emission modes available are upper
side band (USB) voice, lower side band (LSB) voice, amplitude modulation equivalent (AME), or continuous wave
(CW), with a selection of 10, 50, or 175 watts of transmitting power. Transmit tune time is normally less than 1 second. The radio also has automatic link establishment (ALE)
and electronic counter countermeasures (ECCM) frequency
hopping mode. Data messages may be composed and stored
in the receiver/transmitter’s memory. These messages may
be transmitted and received using any operational mode of
the radio set.
b. Power for the radio is provided from the No. 1 dc
primary bus through a circuit breaker marked HF.

3-30

Change 8

FUNCTION

depends

adjacent

Mode switch
MAN

Operating frequency and emission
mode is selected manually. Once
selected, the information is stored
in memory, and can be recalled
using the net selector switch.

PRE

Selects
a
preprogrammed
frequency and emission mode.

ALE

Selects
Automatic
establishment (ALE) mode.

ECCM

Selects
electronic
counter
countermeasure (ECCM) mode.

link

TM 1-1520-237-10

DISPLAY
SCREEN

CURSOR

LINE
SELECT
KEY

FUNCTION
SWITCH

VALUE

SILENT

T/R

STBY

PRE
MAN

ZERO
(PULL)

ALE

ECCM
EMER

BRIGHTNESS
KEY
6

2
1

OFF

KEY

DATA

NET
SELECTOR
SWITCH

VOL
SQL

MODE
SWITCH
AB0988
SA

Figure 3-12.1. Control Display Unit AN/ARC-220

CONTROL/
DISPLAY
EMER
-SQL+ keys

VOL switch

FUNCTION

CONTROL/
DISPLAY

FUNCTION

Selects emergency mode.

OFF

Turns the radio off.

Selects level of squelch from
TONE through 5.
TONE provides no muting or
squelch.
0 gives muting, but no squelch.
1through 5 gives levels of muting
and squelch.
Muting turns off the scanning receiver audio and gives the pilot a
tone when a ALE link is established.

STBY

Turns the radio on, performs bit and
enables fill operations.

SILENT

Prevents the radio from automatically responding to incoming calls
in ALE or ECCM mode. Used during refueling, ordinance loading
and EMCON conditions.

T/R

Allows the radio to transmit and receive in selected operating mode.

ZERO

Erases all loaded data, to include
datafill and keyfill information.

Changes receive audio output level.
VOL settings are displayed for 5
seconds when radio is first powered
up, or when the VOL setting is
changed.

VALUE keys

Increases or decreases a field value
or single character value that is
marked by the cursor.

Function switch

Change 4

3-31

TM 1-1520-237-10

CONTROL/
DISPLAY
Screen displays

FUNCTION
Each line can display up to 20 alphanumeric characters. The 5 characters closest to the line select keys
are used for control selection. See
Table 3-1.1 for advisory messages
and their function.

NOTE
If the radio breaks in and out of squelch,
increase setting as required.
e. -SQL+ switch - Set squelch to 1.
f. Select the desired net (1 through 20), net
selector switch - 1 through +. Use VALUE
keys to select 7 through 20.

3.12A.3 Modes of Operation.
g. ICS Transmitter selector - Position 5.
3.12A.3.1 Manual (MAN) Mode. Use manual mode to
change transmit and receive frequencies, sidebands and
transmit power, and operate the radio manually.
1. To change radio settings:
a. Mode switch - MAN.

h. Radio push-to-talk switch - Press to talk;
release to listen.
3.12A.3.2 Preset (PRE) Mode. Preset mode stores preporgrammed frequencies and emission modes that cannot
be changed by the operator. To use the radio in preset mode,
do the following:

b. Function switch - T/R.
1. Function switch - T/R.
c. Select the desired net (1 through 20), net
selector switch - 1 through +. Use VALUE
keys to select 7 through 20.

2. Mode switch - PRE.
3. -SQL+ switch - Set squelch to 0.

d. EDIT line select key - Press.
NOTE
Changing the receive frequency and mode
will also change the transmission frequency
and mode to the same values. Changing the
transmission frequency and mode will not
change the receive frequency and mode.
e. Edit frequency, emission mode and transmit power by placing the cursor under field
to be edited with CURSOR key, and
change field value with VALUE keys.
f. To end edit and store changed data, RTN
line select key - Press.
2. To operate in manual mode:
a. Function switch - T/R.
b. Mode switch - MAN.
c. -SQL+ switch - Set squelch to 0.
d. VOL switch - Adjust for comfortable listening level.
3-32

Change 4

4. VOL switch - Adjust for comfortable listening
level.
NOTE
If the radio breaks in and out of squelch,
increase setting as required.
5. -SQL+ switch - Set squelch to 1.
6. Select the desired net (1 through 20), net selector switch - 1 through +. Use VALUE keys to
select 7 through 20.
7. ICS Transmitter selector - Position 5.
NOTE
If tune tone is heard, wait until it stops before talking. When radio push-to-talk switch
is pressed, XMT frequency is displayed.
Display returns to preset display when
switch is released.
8. Radio push-to-talk switch - Press to talk; release to listen.

TM 1-1520-237-10

3.12A.3.3 Automatic Link Establishment (ALE)
Mode.

- Press. Time will be transmitted, and radio
will return to scan mode.
2. To receive a ALE call:

WARNING
When in ALE mode, the radio transmits
interrogating signals (sounds) and replies
to ALE calls automatically without operator action. To avoid personnel injury, ensure the function switch is not set to ALE
when personnel are working near the helicopter, during refueling or loading ordinance.

a. INCOMING CALL is displayed, followed by the caller’s ALE address. A short
tone sounds, and LINKED is displayed.
b. ICS Transmitter selector - Position 5.
NOTE
Wait for the calling station to make the first
transmission.

NOTE
Self address must be selected before using
ALE.

c. Radio push-to-talk switch - Press to talk;
release to listen.
3. To place a ALE call:

ALE mode may be used for communications, either normal or link protected, or position reporting.
1. To set up the radio for ALE communications,
do the following:
a. Function switch - T/R.
b. Mode switch - ALE.
c. Select the desired net (1 through 20), net
selector switch - 1 through +. Use VALUE
keys to select 7 through 20.
d. -SQL+ switch - Set squelch to TONE.
e. VOL switch - Adjust for comfortable listening level.
NOTE
Earphone audio is muted until a link is established. If the link is noisy, set squelch to
1. Higher squelch settings are not recommended in this mode.
f. -SQL+ switch - Set squelch to 0.
g. To synchronize time in a link protected
channel, EDIT and SYNC soft keys Press.
h. To broadcast AN/ARC-220 system time as
net control, EDIT, then TXTIM soft keys

a. Select ALE address:
(1) Select the desired net (1 through 20),
net selector switch - 1 through +. To
select 7 through 20, set the selector
switch to the + position and use the
value keys to scroll to the desired selection. Net name and address will be
displayed.
(2) VALUE switch - Press, to scroll
through address list.
(3) If placing an ALE call to an address
not in the list, edit the address as follows: EDIT soft key - Press. Enter address one character at a time with
CURSOR and VALUE switches. To
accept the edit and return to ALE
screen, RTN soft key - Press.
b. ICS Transmitter selector - Position 5.
NOTE
Press ABORT to stop the calling process.
c. Radio push-to-talk switch - Press (and release). CALLING, then LINKED is displayed with a short gong tone in headphone.

Change 8

3-32.1

TM 1-1520-237-10

NOTE
ALE will cancel the link, and return to scan
mode if there is no activity on a link for a
predetermined time as set by the data fill (60
seconds is a typical value). To maintain a
link, press HOLD soft key. When communications are complete, or to return to scan
mode, press SCAN soft key.

under area to change, and VALUE to
change the field to desired value.
e. To save changes and return to top level
screen, RTN soft key - Press.
f. Push-to-talk switch - Press, to tune and
time synchronize the radio.

d. Radio push-to-talk switch - Press to talk;
release to listen.

2. To communicate in ECCM only mode, do the
following:

e. When communication is complete, to return to scanning mode, HOLD, then
SCAN soft key - Press.

a. -SQL+ switch - Set squelch to TONE.

3.12A.3.4 Electronic Counter Countermeasures
(ECCM) Mode. The radio changes frequency in a sequence determined by the ECCM key. Datafill and keyfill
must be loaded prior to using ECCM mode, and system
time must be synchronized between stations. Frequencies
used in hop sets are pretuned in the radio, as ECCM requires frequencies to be changed many times per second.
Frequency hopping is performed in the ECCM mode of
operation. To use this mode, do the following:
1. Initialize the net:
a. Function switch - T/R.

b. VOL switch - Adjust for comfortable listening level.
NOTE
If the frequency is noisy, set squelch to 1.
Higher squelch settings are not recommended in this mode.
c. -SQL+ switch - Set squelch to 0.
d. Press and hold the push-to-talk switch until
XMT READY is displayed. Wait for preamble tones to stop.

b. Mode switch - ECCM.
e. Talk. Release switch to listen.
c. Select the desired net (1 through 12), net
selector switch - 1 through +. Use VALUE
keys to select 7 through 12.
d. To change values on screen, EDIT soft key
- Press. Use CURSOR to position cursor

3-32.2

Change 8

3. Deleted.
3.12A.3.5 Message Mode. The radio can store up to 10
transmit data and 10 received data messages. Each message
may be

TM 1-1520-237-10

500 characters long. Messages are numbered from 1 to 10.
Message 10 is the oldest, and will be deleted if a new
message is received. Messages may be composed using the
AN/ARC-220 CDU dictionary or with a custom dictionary
listing locally generated words, which may be loaded with
datafill.
1. To view a received message:
a. MSG soft key - Press.
b. Use CURSOR keys to scroll left or right,
or up and down in a message.

(2) WORD soft key - Press.
(3) Select the word with VALUE keys.
(4) To insert word with blank in message,
SELECT soft key - Press. If desired,
return to message without inserting a
word by pressing CANCL.
f. To load edited message in R/T memory
and return to top level screen, RTN soft
key - Press.
3. To send a message:

c. Use VALUE keys to page up and down in
a message.
d. To view additional messages, position cursor under message number with CURSOR
keys. Use VALUE keys to scroll to the
next message number.

a. Access PRGM MSG screen by pressing
MSG, then PRGM soft keys.
b. Select message to send as desired by placing cursor under message number, and
pressing VALUE keys until desired message is displayed.

e. To retain received messages, RTN soft key
- Press.
f. To delete received messages, position the
cursor under the message number and DEL
soft key - Press, until messages are deleted.
To return to top screen, RTN soft key Press.
2. To edit or compose a message:
a. MSG soft key - Press.
b. From MESSAGE screen, PGRM soft key
- Press.
c. Select message to be edited by placing cursor under the message number with CURSOR keys, and change number with
VALUE keys.
d. Edit message by placing cursor under area
to be changed. Use VALUE keys to
change one character at a time. Press DEL
to delete one character at a time.

NOTE
Message will be sent to currently selected
address (ALE modes) or transmitted on the
currently selected frequency and mode
(MAN, PRE, or ECCM).
c. SEND soft key - Press.
3.12A.4 Operation.
3.12A.4.1 Starting Procedure.
1. Function switch - STBY. SYSTEM TESTING
is displayed while power up built in test (PBIT)
is in process. SYSTEM - GO will be displayed
upon successful completion of PBIT.
2. FILL line select key - Press. Status of PRE,
ALE, ECCM, and EMER modes will be displayed.

e. To insert a word from the dictionary in a
message do the following:

3.12A.4.2 Load Presets. Datafill contains preset frequencies, scan lists, addresses, data messages, and non secure information needed for ALE/ECCM operation. If the
DTD is configured to receive data, it may be copied from
the radio to the DTD by pressing COPY line select key on
the DATA FILL page.

(1) Position cursor where the word is to
be inserted

1. Initialize the data transfer device (DTD). Connect the DTD to the DATA connector.

Change 4

3-32.3

TM 1-1520-237-10

2. ZERO line select key - Press.

2. With the FILL page selected, DATA line select key - Press.

3. Select key to zero with VALUE keys. Default
is all keys.

NOTE
Pressing RTN line select key on DATA
FILL page stops the fill process.

NOTE

3. On the DATA FILL page, FILL line select
key - Press. FILL ENABLED screen will appear.

If you do not want to zero the key, press
NO. The FILL screen will then appear.
4. Confirm zero by pressing YES line select key.
ZEROIZE advisory message will appear, followed by the FILL screen.

4. Start data fill on DTD. Monitor DTD to see
when data transfer is complete.
3.12A.4.3 Load Secure Keys. Key fill contains secure
information needed for ALE link protection and ECCM
operation.
1. Initialize the DTD. Connect the DTD to the
KEY connector.

3.12A.4.5 Emergency
(EMER)
Operation. The
mode, frequency, and net to be used in the EMER position
is determined by the datafill. To use the emergency mode,
do the following:
1. Function switch - T/R.

2. With the FILL page selected, KEY line select
key - Press.

2. Mode switch - EMER.
3. ICS Transmitter selector - Position 5.

NOTE

4. Radio push-to-talk switch - Press to talk; release to listen.

Pressing RTN line select key on KEY FILL
page stops the fill process.
3. On the KEY FILL page, LOAD line select key
- Press. FILL ENABLED message will appear.

3.12A.5 Shutdown.
1. Function switch - OFF.

4. Start keyfill on DTD. Monitor DTD to see when
data transfer is complete.

2. To erase all preprogrammed information, Function switch - Pull and turn to ZERO (PULL).

3.12A.4.4 Zero Secure Keys.
1. Access KEY FILL page. From FILL screen,
KEY fixed function key - Press.

3.12A.6 Messages. Table 3-2 lists display advisory
messages that may appear during operation of the radio:

Table 3-2. AN/ARC-220 Messages

ADVISORY

MEANING

ACTION

ALE - NO DATA

ALE mission data not loaded.

Load mission data.

ALE - NO KEYS

ALE link protection keys not loaded.

Load keys.

CALL FAIL

Radio failed to complete an outgoing call.

CALLING

Radio is placing an ALE call to another address.

3-32.4

Change 8

TM 1-1520-237-10

Table 3-2. AN/ARC-220 Messages (Cont)

ADVISORY

MEANING

ACTION

CDU FAIL

Radio set control is inoperative.

CHANNEL BUSY

ALE or ECCM net is in use.

CHANNEL INOP

ALE or ECCM keys are not loaded, or not correct.

CHECK MSG

A data message has been received.

COMPLETE

Indicated power-up BIT is complete.

COPY COMPLETE

Copying process finished successfully.

COPY FAIL

Copying process was unsuccessful.

COPYING DATA

The radio is copying datafill contents from DTS.

ECCM - NO DATA

ECCM data not installed.

Load mission data.

ECCM - NO KEYS

ECCM keys not installed.

Load keys.

EMER

Mode or net selected for emergency communication is inoperative.

EMERG - NO KEYS

No keys available for net selected for emergency
communication.

EOM

End of message.

EXT FAIL

Radio failed due to external device, such as antenna.

GO DATA

Link quality analysis values too low for reliable
voice communication; data transmissions recommended.

GPS FAIL

Position report could not be issued.

GPS TIME FAIL

Current time could not be established via GPS receiver.

HELD

ALE call being held in specific frequency by operator.

INCOMING CALL

Another radio is establishing an ALE link.

INOP MODES EXIST

Warning to expect inoperative modes.

LINKED

An ALE link is established.

Wait or try another net.

Load keys.

Change 8

3-32.5

TM 1-1520-237-10

Table 3-2. AN/ARC-220 Messages (Cont)

ADVISORY

MEANING

LOAD COMPLETE

Keys and data successfully loaded into radio.

LOAD FAIL

Keys and data not successfully loaded into radio.

LOADING DATA

Radio currently loading data.

LOADING KEYS

Radio currently loading keys.

MSG ABORT

Radio discontinuing sending of current message.

NET INOP

Selected net contains no data, corrupted data, or
hardware cannot support the selected mode of operation.

NO AUTO XMT

Radio has been instructed not to make any automatic transmissions.

NO DATA

Database is not filled with necessary data to perform requested operations.

NO KEYS LOADED

Keys are not loaded for current selected mode or
net.

NO RCVD MSGS

No messages have been received.

PAC FAIL

Failure of radio in PA coupler.

PLGR

Precision lightweight GPS receiver.

POSN RPT FAIL

Current GPS position report not loaded.

PRE - NO DATA

Preset data not loaded.

PTT FOR XMIT BIT

Instruction to press microphone PTT switch to enable transmission BIT.

RCV BIT - GO

Receiver BIT functions completed without failure.

RCV READY

Ready to receive ECCM transmissions.

RCVG PREAMBLE

ECCM preamble being received.

RCVG DATA

Radio currently receiving data.

RT-CDU COMM FAIL

Receiver-transmitter is failing to communicate
with the radio set control.

RT FAIL

Receiver Transmitter inoperative.

3-32.6

Change 8

ACTION

TM 1-1520-237-10

Table 3-2. AN/ARC-220 Messages (Cont)

ADVISORY

MEANING

ACTION

RX-TX DEGRADED

Receive and transmit capabilities are degraded.

RX-TX FAIL

Radio cannot receive or transmit.

SENDING DATA

Radio currently sending data.

SENDING POSN

Sending GPS position report.

SOUND

Radio sending an ALE sound.

SYNCING

Time synchronization being performed.

TESTING

BIT in progress.

TIME SYNC FAIL

Radio failed in attempt to synchronize.

TRANSEC FAIL

BIT detected a failure that will not allow ECCM
operation.

TUNE XX%

Indicates percentage of ECCM frequencies tuned
for current net.

TUNING

Radio is currently tuning itself.

TX DEGRADED

BIT detected a failure that is causing transmission
capability to be degraded.

TX FAIL

Radio cannot transmit.

UNSYNC

ECCM is not synchronized.

UNTUNED

An ECCM hop set is not tuned.

XMT READY

Radio is ready to transmit in ECCM mode.

ZEROIZED

All mission datafill and keys have been erased.

3.12B TSEC/KY-100 SECURE COMMUNICATION
SYSTEM.
The TSEC/KY-100 provides secure, half duplex voice,
digital data, analog data and remote keying capabilities for
transmission over the AN/ARC-220 HF radio. It has six
operational modes, and can store often used settings on
presets. Power is supplied from the No. 1 dc primary bus
through a circuit breaker marked HF SCTY SET.
3.12B.1 Controls and Functions. The KY-100 is controlled by a control display unit (CDU) located behind the

lower console (Figure 3-12.2). The function of each control
and display is as follows:
CONTROL

FUNCTION

AUDIO

Speaker for audio tones.

CIK

Cryptographic Ignition Key. Not
used in this installation.

FILL connector

Used to connect external fill device
to KY-100.

Change 8

3-32.7

TM 1-1520-237-10

CONTROL

FILL
DSPL
OFF

KY−
100
CIK

CT
PT
MODE
1

2 3

MAN
PWR
OFF

6
PRESET

OFL

Sets KY-100 to off line mode.
Disables communications and
accesses screens to select mode
settings, test, and fill screens.

Select emergency back-up key.
ALL

Erases all cryptographic data (keys)
except the emergency back-up key.

BAT
5

3.12B.2 Modes of Operation.

FUNCTION
Function keys used to access and
navigate in software menus.

DSPL OFF

Varies light intensity of display.
Display turned off in OFF position.
Varies light intensity of backlit
display panel. Display turned off in
OFF position.

PRESET switch

Controls power to set, and which
key is active.

PWR OFF

Removes power from set.

MAN

Manual rekeying enabled.

1,2,3,4,5,6

Selects preset settings for use.

REM

Allows control of KY-100 from a
remote control unit (RCU).

Change 9

Allows
cooperative
terminal
rekeying in receive mode.

Z
(PULL)

INIT, g and d

3-32.8

REM

Figure 3-12.2. KY-100 Secure Communication
Control Panel

MODE switch

Sets KY-100 to ciphertext mode.

PNL OFF

BRT

AB0987

CONTROL

Sets KY-100 to plaintext mode.

U
4

PNL
OFF

INIT

OFL EB Z ALL
(PULL)
RK

AUDIO

FUNCTION

3.12B.2.1 Plaintext (PT) Mode. The ICS voice signal
is routed through the KY-100 to and from the HF radio,
with no processing. Radio transmits and receives unencrypted information.
3.12B.2.2 Cyphertext (CT) Mode. The ICS voice signal is routed to the KY-100, where it is processed, encrypted and sent to the HF radio for transmission. Received
audio signals from the HF radio are processed, decoded,
and sent to ICS. Unencrypted information is routed through
the KY-100 if CT ONLY is not selected in configuration
settings. Non cooperative rekey receive is possible only in
this mode.
3.12B.2.3 Rekey Mode. Use this mode to fill crypto information. The data transfer device must be connected to
FILL to load keys.
3.12B.2.4 Off-line (OFL) Mode. For maintenance use
to configure and test the system. Communications are not
possible in this mode.
3.12B.2.5 Emergency Backup (EB) Mode. Enables a
zeroized terminal to be used for voice privacy operation,
only. Key is not erased when terminal is zeroized. This
mode is not to be used to transmit classified information.
3.12B.2.6 Zeroize (Z ALL) Mode. Erases all keys in
the KY-100 except the emergency back-up key.

TM 1-1520-237-10

3.12B.3 Operation.
3.12B.3.1 Keyfill Operation. When there are no TEKs
in the KY-100 at start up, the display will read CLd STRT.
If there are TEKs in the terminal, skip steps 3 and 6, and
load or update keys as required.

3.12B.3.2 Normal Operation.
1. MODE switch - PT, or CT.
2. PRESET switch - MAN, 1, 2, or 3.
3.12B.3.3 Emergency Operation.

1. MODE switch - OFFLINE.
NOTE
2. PRESET switch - MAN.
3. Wait until CLd STRT is displayed, then INIT
key - Press.
4. Connect a fill device to FILL connector.
5. Turn on device and select key to be loaded.
6. INIT key - Press. At the end of the fill sequence, a tone should be heard in the headset,
and KEY 1 01, CIK OK, and PASS will appear. The key that was loaded is stored in fill
position 1.
7. To fill the rest of the keys, push the d or g key
until KEY OPS is displayed.
8. INIT key - Push twice. LOAD KEY, then
LOAD X will be displayed with the flashing X
being the number of currently selected key location.
9. Press the d or g key until the desired location
(1, 2, 3, 4, 5, 6, or U) is displayed.

Emergency key is not secure. Do not transmit classified information in this mode.
1. MODE switch - EB.
2. PRESET switch - MAN, 1, 2, or 3.
3.12B.3.4 Zeroize All Keys.
NOTE
Power does not have to be applied to unit to
zero all keys.
Emergency backup key is not zeroized in
this procedure.
1. MODE switch - Pull and rotate to Z ALL
(PULL).
3.12B.3.5 Zeroize Specific Keys.
1. MODE switch - OFFLINE.

10. INIT key - Press. The entire LOAD X display
will flash.

2. UP ARROW, or RIGHT ARROW soft key Press, until KEY OPS is displayed.

11. Turn on device and select key to be loaded.

3. INIT key - Press. LOAD KEY will be displayed.

12. INIT key - Press. At the end of the fill sequence, a tone should be heard in the headset,
and KEY X will appear. The display will then
change to LOAD X with the flashing X being
the number of currently selected key location.
13. Repeat steps 9. through 11. until all required
locations are filled.
14. When all keys are transferred, turn off the fill
device, and disconnect it from FILL connector.
15. To exit key load, place MODE switch out of
OFFLINE.

4. UP ARROW, or RIGHT ARROW key Press, until ZERO is displayed.
5. INIT KEY soft key - Press. ZERO X, with a
flashing number (X) appears. The flashing number indicates the currently selected key to be
zeroized.
NOTE
Number of keys is 1 through 6 TEKs, U
(used to update internal keys) and Eb (emergency backup key).

Change 8

3-32.9

TM 1-1520-237-10

6. UP ARROW, or RIGHT ARROW key Press, until key number to zeroize is displayed.

9. Repeat steps 6 through 8 to zero other key positions, as desired.

7. INIT key - Press. The entire ZERO X will
now flash.

10. When all desired key positions are zeroized,
MODE switch - Move to any other position.

8. INIT key - Press. The screen will blank while
zeroizing process takes place. When zeroizing
is complete, a tone will be heard in the headset,
the display will briefly change to ZEROED X,
and then revert to ZERO X.

3.12B.4 Shutdown.
PRESET switch - PWR OFF.
3.12B.5 Configuration.
Configure as described in Appendix C.

3-32.10

Change 8

TM 1-1520-237-10

Section III NAVIGATION
CONTROL

KILOCYCLES
CW

A
D
F
R
C
V
R

T
U
N
E

AUDIO

VOICE

Mode selector
switch

TEST

COMP
OFF

ANT

FUNCTION

LOOP

OFF

Turns power off.

COMP

Provides operation as an ADF.

ANT

Provides for operation as an AM
receiver using sense antenna.

LOOP

Provides for receiver operation as a
manual direction finder using loop
only.

LOOP
L

100 KILOHERTZ COARSE
TUNE CONTROL

10 KILOHERTZ FINE
TUNE CONTROL
MODE SELECTOR

AA0360
SA

Figure 3-13. LF/ADF Control Panel
C-7932/ARN-89

LOOP
control
switch

3.13 DIRECTION FINDER SET AN/ARN-89. (LF/
ADF).

AUDIO

Adjusts volume.

100 Kilohertz
coarse-tune
control knob

Tunes receiver in 100-kHz steps as
indicated by first two digits of
KILOCYCLES indicator.

10 Kilohertz fine
-tune
control
knob

Tunes receiver in 10 kHz steps as
indicated by last two digits of
KILOCYCLES indicator.

Direction Finder set AN/ARN-89 (Figure 3-13) is an
airborne, low frequency (LF), automatic direction finder
(ADF) radio, that provides an automatic or manual compass bearing on any radio signal within the frequency range
of 100 to 3,000 kHz. The ADF can identify keyed or continuous wave (CW) stations. The ADF displays the bearing
of the helicopter relative to a selected radio transmission on
the horizontal situation indicator No. 2 bearing pointer
(Figure 3-31 ). When ADF is selected on the MODE SEL
panel (Figure 3-32) three modes of operation permit the
system to function: as a CW automatic direction finder, as a
CW manual direction finder or as an amplitude-modulated
(AM) broadcast receiver. Power to operate the Direction
Finder AN/ARN-89 is provided by No. 1 dc primary bus
through a circuit breaker, marked ADF, and the ac essential
bus through a circuit breaker, marked 26VAC INST.
3.13.1 Antennas. The ADF sense antenna is a part of
the VHF/FM No. 2, VHF/AM, antenna (Figure 3-1) under
the nose section of the helicopter. The ADF loop antenna is
flush-mounted, under the center fuselage section.

L-R

Provides manual left and right
control of loop when operating
mode selector in LOOP position. It
is spring loaded to return to center.

CW, VOICE,
TEST switch
CW (COMP
mode)

Enables tone oscillator to provide
audible tone for tuning to CW
station, when mode function switch
is at COMP.

CW (ANT or
LOOP mode)

Enables beat frequency oscillator to
permit tuning to CW station, when
mode function switch is at ANT or
LOOP.

VOICE

Permits low frequency receiver to
operate as a receiver with mode
switch in any position.

3.13.2 Controls and Functions. Controls for the LF/
ADF receiver are on the front panel of the unit (Figure
3-13). The function of each control is as follows:

Change 8

3-32.11

TM 1-1520-237-10

CONTROL

FUNCTION

4. ICS NAV switch - ON.
5. To test the ADF, when required:

TEST (COMP
mode)

TUNE meter

Provides slewing of loop through
180° to check operation of receiver
in COMP mode. (Switch position
is inoperative in LOOP and ANT
mode.)
Indicates relative signal strength
while tuning receiver to a specific
radio signal.

KILOCYCLES
indicator

Indicates operating frequency to
which receiver is tuned.

a. CW, VOICE, TEST switch - TEST.
Check to see that No. 2 bearing pointer
changes about 180°.
b. CW, VOICE, TEST switch - Release.
3.13.3.4 LOOP Mode Operation. Manual direction
finding uses the LOOP mode.
1. Mode selector switch - LOOP.

3.13.3 Operation.

2. ICS NAV switch - ON.

3.13.3.1 Starting Procedure.

3. Turn LOOP L-R switch to L (left) or R (right)
to obtain an audio null and a TUNE indicator
null. Watch HSI No. 2 bearing pointer for a
display of magnetic bearing to or from ground
station as read against the compass card. In this
mode of operation, two null positions 180°
apart are possible.

1. ICS NAV receiver selector - ON.
2. Mode selector - COMP, ANT, or LOOP.
3. Frequency - Select.
4. CW, VOICE, TEST switch - CW or VOICE
as appropriate.
5. ICS NAV switch - ON.
6. Fine tune control - Adjust for maximum upward indication on TUNE meter.
7. AUDIO control - Adjust as desired.
3.13.3.2 ANT Mode Operation.
1. Mode selector - ANT.
2. ICS NAV switch - ON.
3. Monitor receiver by listening.
3.13.3.3 COMP Mode Operation.
1. Mode selector - COMP.
2. MODE SEL BRG 2 HSI/VSI switch - ADF.
3. The horizontal situation indicator No. 2 bearing
pointer displays the magnetic bearing to the
ground station from the helicopter, as read
against the compass card, when ADF is selected on the MODE SEL BRG 2 switch.
3-32.12

Change 8

3.13.4 Stopping Procedure. Mode selector - OFF.
3.14 DIRECTION FINDER SET AN/ARN -149 (LF/
ADF) (IF INSTALLED).
The AN/ARN -149 (Figure 3-14) is a low frequency
(LF), automatic direction finder (ADF) radio, providing
compass bearing capability within the frequency range of
100 to 2199.5 kHz. The ADF has two functional modes of
operation: ANT and ADF. The antenna (ANT) mode functions as an aural receiver, providing only an aural output of
the received signal. The ADF mode functions as an automatic direction finder, providing a relative bearing-tostation signal to the horizontal situation indicator No. 2
bearing pointer and an aural output. A TONE submode of
operation can be selected in either ANT or ADF mode,
providing a 1000-Hz aural output to identify keyed CW
signals. Power is provided to the LF/ADF system by the
No. 1 dc primary bus through a circuit breaker, labeled
ADF, and the ac essential bus through a circuit breaker,
labeled 26 VAC INST.
3.14.1 Antennas. The antenna system is a single combination antenna containing both loop and sense elements.
The RF signal from one loop element is modulated with a
reference sine signal while the other loop element is modulated with a reference cosine signal. The two modulated
signals are combined, phase shifted 90°, and amplified. The
resulting loop signal is summed with the sense antenna sig-

TM 1-1520-237-10

nal and sent to the ADF radio for visual and aural execu-

tion. The antenna configuration is flush mounted under the
bottom cabin fuselage (Figure 3-1).

Change 8

3-32.13/(3-32.14 Blank)

TM 1-1520-237-10

FREQUENCY
CONTROLS AND
INDICATORS

A
D
F
MAN

TEST

TAKE
CMD

VOL

ADF

2182
500

MANUAL
2182 / 500
SELECT

ANT
OFF

TONE

TEST / (OFF) / TONE
SELECT

VOLUME
ADJUST

TAKE
COMMAND
SELECT

ADF / ANT / OFF
SELECT
FS0015A
SA

Figure 3-14. LF/ADF Control Panel AN/ARN-149
3.14.2 Controls and Functions. Controls and frequency digit displays are on the front of the ADF control
panel (Figure 3-14). The function of each control is as follows:

CONTROL
Frequency controls
and indicators

CONTROL
TEST/(OFF)/
TONE select

TEST position (up) is a momentary
position that enables a self-test.
Center position is off. TONE
position (down) enables the tone
generator for CW operation.

VOL adjust

A 12-position switch controlling
volume in 12 discrete steps.

TAKE CMD select

Used in a dual ADF control panel
installation allowing each to take
control of the receiver away from
the other. Not used in this
installation.

FUNCTION
Controls and indicates the selected
frequency when MAN/2182/500
switch is in MAN.

MAN/2182/500
select
MAN

Enables the frequency controls and
indicators.

2182

Selects 2182 kHz as the operating
frequency.

500

Selects 500 kHz as the operating
frequency.

FUNCTION

ADF/ANT/OFF
select
ADF

Applies power to system and turns
on ADF and aural capability.

ANT

Applies power to system and turns
on antenna or aural function only.

OFF

Removes power from system.

Change 1

3-33

TM 1-1520-237-10

3.14.3 Operation.

3.14.4 Stopping Procedure. ADF/ANT/OFF switch OFF.

3.14.3.1 ANT (Aural Only) Operation.
1. ICS NAV reciever selector switch - ON.
2. ADF/ANT/OFF switch - ANT.
3. MAN/2182/500 switch - As desired.
If MAN is selected in step 3:
4. Frequency switches - Select.
5. VOL control - Adjust as desired.
6. TEST/(OFF)TONE switch - TONE.
3.14.3.2 ADF Operation.
1. ICS NAV receiver selector switch - ON.
2. HSI/VSI MODE SEL BRG 2 switch - ADF.
3. ADF/ANT/OFF switch - ADF.
4. MAN/2182/500 switch - As desired.
If MAN is selected in step 4:

3.15 RADIO RECEIVING
(VOR/ILS/MB).

SET

AN/ARN-123(V)

Radio set AN/ARN 123(V) (Figure 3-15) is a very highfrequency receiver that operates from 108.00 to 117.95
MHz. Course information is presented by the VSI course
deviation pointer and the selectable No. 2 bearing pointer
on the horizontal situation indicator. The combination of
the glide slope capability and the localizer capability makes
up the instrument landing system (ILS). The marker beacon
portion of the receiver visually indicates on the VSI MB
advisory light, and aurally indicates on the headphones, of
passage of the helicopter over a marker beacon transmitter.
The receiving set may be used as a VOR receiver, or ILS
receiver. The desired type of operation is selected by tuning
the receiving set to the frequency corresponding to that
operation. ILS operation is selected by tuning to the odd
tenth MHz frequencies between 108.0 and 112.0 MHz.
VOR operation is selected by tuning in .050 MHz units to
the frequencies between 108.0 and 117.95 MHz, except the
odd tenth MHz between 108.0 and 112.0 MHz, which are
reserved for ILS operation. The three receiver sections do
the intended functions independent of each other. Performance degradation within any one of the major sections
will not affect the performance of the others. Power for the
AN/ARN-123 is provided from the dc essential bus through
a circuit breaker, marked VOR/ILS.

5. Frequency controls - Select.
6. VOL control - Adjust as desired.
If CW operation is desired:
7. TEST/(OFF)/TONE switch - TONE.
8. Verify horizontal situation indicator (HSI) No.
2 bearing pointer displays appropriate relative
bearing-to-the-station.
If self-test is required:
9. TEST/(OFF)/TONE switch - TEST (position
up and hold).
10. No. 2 bearing pointer deflects 90° away from
original reading.
11. TEST/(OFF)/TONE switch - Release to off.
12. Verify No. 2 bearing pointer returns to original
reading.

3-34

Change 10

NOTE
Tuning to a localizer frequency will automatically tune to a glide slope frequency,
when available.
3.15.1 Antenna. The VOR/LOC antenna system (Figure
3-1), consists of two blade type collector elements, one on
each side of the fuselage tail cone. The glide slope antenna
is mounted under the avionics compartment in the nose.
The antenna provides the glide slope receiver with a
matched forward-looking receiving antenna. The marker
beacon antenna is flush-mounted under the center section of
the fuselage.
3.15.2 Controls and Functions. The controls for the
VOR/ILS/MB receivers are on the front panel of the unit.
The function of each control is as follows:

TM 1-1520-237-10

2. NAV VOL OFF control - On.
3. Frequency - Select.
NAV VOL

MB VOL

4. MODE SEL BRG 2 switch - VOR.

108.00
OFF

OFF
VOR / MB
TEST

5. MODE SEL VOR/ILS switch - VOR.

MB SENS

3.15.3.2 VOR/Marker Beacon Test.

HI
LO

MEGAHERTZ
TUNE CONTROL

FREQUENCY
INDICATOR

HUNDRETHS
MEGAHERTZ
TUNE CONTROL

NOTE
If acceptable signal is not received, test will
not be valid.
AA0526
SA

Figure 3-15. Radio Receiving Set
AN/ARN-123(V)

CONTROL

FUNCTION

1. HSI CRS set 315° on COURSE set display,
pilot and copilot.
2. VOR/MB TEST switch - Down and hold. The
MB light on the VSI should go on.
3. HSI VOR/LOC course bar and VSI course deviation pointer - Centered 6 1 dot.

NAV VOL-OFF
control

Turns VOR/ILS receiver on and
off, adjusts volume.

MB VOL-OFF
control

Turns marker beacon receiver on
and off; adjusts volume.

Megahertz tune
control

Tunes VOR/ILS receiver in MHz
as indicated on frequency indicator.

Hundredths
megahertz tune
control

Tunes VOR/ILS receiver in
hundredths MHz as indicated on
frequency indicator.

3.15.3.4 ILS (LOC/GS) Operation. ILS operation frequency - Set.

VOR/MB TEST
control

Activates VOR test circuit and MB
receiver lamp self-test circuits.

3.15.3.5 Marker Beacon (MB) Operation.

MB SENS HILO control

For controlling MB sensitivity.

4. No. 2 bearing pointer should go to the 310° to
320° position.
5. To-from arrow should indicate - TO.
6. VOR/MB TEST switch - Release.
3.15.3.3 VOR Operation. Course - Set.

1. MB VOL OFF switch - On.
2. MB SENS switch - As desired.

Decreases receiver sensitivity by
shortening time transmitted signal
will be received.

3.15.3.6 VOR Communications Receiving Operation. Frequency - Set.

Increases receiver sensitivity by
lengthening time transmitted signal
will be received.

3.15.4 Stopping Procedure. NAV VOL OFF switch OFF.

3.15.3 Operation.
3.15.3.1 Starting Procedure.
1. ICS AUX selector - ON.

3.16 RADIO RECEIVING SET AN/ARN-147(V)(VOR/
ILS/MB)(IF INSTALLED).
Radio set AN/ARN-147 (V) (Figure 3-16) is a very high
frequency receiver, capable of operating from 108.0 to
126.95 MHz. Course information is presented by the vertical situation indicator deviation pointer and the selectable

Change 9

3-35

TM 1-1520-237-10

No. 2 bearing pointer on the horizontal situation indicator.
The combination of the glide slope and localizer capabilities makes up the instrument landing system (ILS). The
marker beacon portion of the receiver visually indicates on
the vertical situation indicator MB advisory light, and aurally signals over the headphones helicopter passage over a
transmitting marker beacon. The radio set may be used as a
VHF omnirange (VOR) or ILS receiver. The desired type
of operation is selected by tuning the receiving set to the
frequency corresponding to that operation. ILS operation is
selected by tuning to the odd tenth MHz frequencies from
108.0 to 111.95 MHz. VOR operation is selected by tuning
from 108.0 to 126.95 MHz, except the odd tenth MHz from
108.0 to 111.95 MHz reserved for ILS operation. The three
receiver sections do the intended functions independent of
each other. Performance degradation within any one of the
major sections will not affect performance of the others.
Power for the AN/ARN-147 is provided from the dc essential bus through a circuit breaker, labeled VOR/ILS.

FUNCTION

MB HI/LO select

Varies marker beacon
sensitivity (high or low).

(MB)

MHz digits select

Changes frequency in 1-MHz steps
over the range of control (first three
digits).

MB VOL adjust

Varies marker beacon (MB) audio
gain of the associated receiver.

3.16.3 Operation.
3.16.3.1 Starting Procedure.
1. ICS AUX receiver selector switch - ON.
2. TEST/(pwr) ON/OFF switch - ON (center position).

NOTE
Tuning to a localizer frequency will automatically tune to a glide slope frequency
when available.
3.16.1 Antennas. The VOR/LOC antenna system (Figure 3-1) consists of two blade type collector elements, one
on each side of the fuselage tail cone. The glide slope antenna is mounted under the avionics compartment in the
nose. The antenna provides the glide slope receiver with a
matched forward looking receiving antenna. The marker
beacon antenna is flush-mounted under the center section of
the fuselage.
3.16.2 Controls and Functions. The controls for the
VOR/ILS/MB receivers are on the front of the control panel
(Figure 3-16). The function of each control is as follows:
CONTROL

CONTROL

4. KHz (last two digits) control - Select.
5. NAV VOL control - Adjust.
6. MODE SEL BRG 2 switch - VOR.
7. MODE SEL VOR/ILS switch - VOR.
8. CIS MODE SEL NAV/ON switch - As desired.
3.16.3.2 VOR/Marker Beacon Test.
NOTE
Test will not be valid if signal reception is
invalid.

FUNCTION

Digit window

Indicates
frequency.

operating

1. HSI CRS control (pilot and copilot) - Set 315°
in course display.

NAV VOL adjust

Varies navigation (VOR/LOC)
audio gain of associated receiver.

2. TEST/(pwr) ON/OFF switch - TEST (position
up and hold). VSI MB advisory light goes on.

KHz digits select

Changes frequency in 50-kHz steps
over the range of control (last two
digits).

3. HSI VOR/LOC course bar and VSI course deviator pointer - Centered (61 dot).

TEST/(power) ON/
OFF select

3-36

selected

3. MHz (first three digits) control - Select.

Controls application of power to
the associated receiver. Controls
VOR/marker beacon test.

Change 9

4. No. 2 bearing pointer - 315° (65°).
5. To-from arrow - TO.

TM 1-1520-237-10

DIGIT
WINDOW

MARKER BEACON
VOLUME ADJUST

N
A
V

NAV VOLUME
ADJUST

MB
VOL

NAV
VOL

MB
HI
LO

TEST
ON
OFF

MHZ
DIGITS
SELECT

KHZ DIGITS
SELECT

MARKER BEACON
HI / LO SELECT

TEST / (POWER)
ON / OFF SELECT

FS0016A
SA

Figure 3-16. VOR/ILS/MB Control Panel AN/ARN-147 (V)
6. TEST/(pwr) ON/OFF switch - Release.
3.16.3.3 VOR Operation. HSI CRS control - Course select.
3.16.3.4 ILS (LOC/GS) Operation.
1. ILS operation frequency/volume - Set.
2. HSI CRS control - Course select.
3. CIS MODE SEL NAV/ON switch - As desired.
3.16.3.5 Marker Beacon (MB) Operation.
1. ICS NAV receiver selector switch - ON.
2. MB HI/LO switch - As desired.
3. MB VOL control - Adjust as desired.
3.16.3.6 VOR Communications Receiving Operation. Frequency/Volume - Set.

3.17 DOPPLER NAVIGATION SET
AN/ASN-128. UH
The Doppler navigation set, AN/ASN-128, in conjunction with the helicopter’s heading and vertical reference
systems, provides helicopter velocity, position, and steering
information from ground level to 10,000 feet. To achieve
best results with the set, pitch and roll angles should be
limited to 30° pitch and 45° roll, and moderate maneuver
rates should be employed. The Doppler navigation system
is a completely self-contained navigation system and does
not require any ground-based aids. The system provides
world-wide navigation, with position readout available in
both Universal Transverse Mercator (UTM) and Latitude
and Longitude (LAT/LONG)(Figure 3-22). Navigation and
steering is done using LAT/LONG coordinates, and a bilateral UTM-LAT/LONG conversion routine is provided for
UTM operation. Up to ten destinations may be entered in
either format and not necessarily the same format. Present
position data entry format is also optional and independent
of destination format. Power to operate the AN/ASN-128 is
provided from No. 1 dc primary bus through a circuit
breaker marked DPLR, and from the ac essential bus
through a circuit breaker, marked 26 VAC DPLR, refer to
TM 11-5841-281-12.

3.16.4 Stopping Procedure. TEST/(pwr) ON/OFF
switch - OFF.

Change 1

3-37

TM 1-1520-237-10

LEFT
DISPLAY
LAMPS

CENTER
DISPLAY
LAMPS

RIGHT
DISPLAY
LAMPS

DIM
DIST / BRG
TIME

PP
GS
D
P
L
R

TK
XTK
TKE

DEST
TGT

WIND

DISPLAY

TEST

FLY−TO
DEST

UTM
LAT /
LONG

LAMP
TEST

LEFT

MID

RIGHT

ABC
1

DEF
2

GHI
3

JKL
4

MNO
5

PQR
6

STU
7

VWX
8

YZ
9

CLR

ENT

BACKUP

Places navigation set in estimated
mode of operation or estimated
velocity mode of operation.

WIND
DIR

Selects navigation data for display.
SP/

XTK/TKE
(Left Display)
KEYBOARD

3
BACK
UP

OFF

Select
latitude/longitude
navigational mode of operation.

(Right Display)

MODE

AA0663
SA

Figure 3-17. Doppler Navigation Set
AN/ASN-128
3.17.1 Antenna. The Doppler antenna (Figure 3-1) consists of a combined antenna/radome and a receivertransmitter housing below copilot’s seat. The combination
antenna/radome uses a printed-grid antenna.
3.17.2 Controls, Displays, and Function. The control and displays for the Doppler are on the front panel
(Figure 3-17). The function of each control is as follows:

CONTROL/
INDICATOR
MODE selector

FUNCTION
Selects Doppler Navigation Mode
of operation.

OFF

Turns navigation set off.

LAMP TEST

Checks operation of all lamps.

TEST

Initiates built-in-test exercise for
navigation set.

UTM

Selects
Universal
Mercator (UTM)
mode of operation.

3-38

Not applicable.

SPH
VAR

SP / DIR

N
A

ALPHA
DEST
DISP

FUNCTION

LAT/LONG

DISPLAY
selector

TGT
STR

KYBD

MEM MAL

CONTROL/
INDICATOR

TARGET
STORAGE
INDICATOR

Transverse
navigational

GS-TK
(Left Display)
(Right Display)
PP with switch
set to UTM
(Center Display)

Distance crosstrack (XTK) of
initial course to destination in km
and tenths of a km.
Track angle error (TKE) in degrees
displayed as right or left of bearing
to destination.
Ground speed (GS) in km/hr.
Track angle (TK) in degrees
TRUE.
Present position UTM zone.

(Left Display)

Present position UTM area square
designator and easting in km to
nearest ten meters.

(Right Display)

Present position UTM area
northing in km to nearest ten
meters.

PP with MODE
switch set to
LAT/LONG
(Left Display)

Present position latitude in degrees,
minutes and tenths of minutes.

(Right Display)

Present position longitude in
degrees, minutes and tenths of
minutes.

DIST/BRGTIME
(Center
Display)
(Left Display)

Time to destination selected by
FLY TO DEST (in minutes and
tenth of minutes).
Distance to destination selected by
FLY TO DEST (in km and tenths
of a km).

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

(Right Display)

Bearing to destination selected by
FLY TO DEST (in degrees MAGNETIC).

DEST-TGT
(Mode switch set
to UTM)
(Center
Display)

UTM zone of destination selected
by DEST DISP thumbwheel.

(Left Display)

UTM area and easting of destination set on DEST DISP thumbwheel.

(Right Display)

Northing of destination set on
DEST DISP thumbwheel.

DEST-TGT
(Mode switch set
to LAT/LONG
(Left Display)

Latitude (N 84° or S 80° max.) of
destination set on DEST DISP
thumbwheel.

(Right Display)

Longitude of destination set on
DEST DISP thumbwheel.

SPH-VAR
(Left Display)
(Right Display)

CONTROL/
INDICATOR
KYBD pushbutton

Used in conjunction with the keyboard to allow data to be displayed
and subsequently entered into the
computer when the ENT key is
pressed.

DEST DISP
thumbwheel
switch

Destination display thumbwheel
switch is used along with DESTTGT and SPH-VAR position of
DISPLAY switch to select destination whose coordinates or magnetic
variation are to be displayed, or to
be entered. Destinations are 0
through 9, P (Present Position) and
H (Home).

Keyboard

Used to set up data for entry into
memory. When the DISPLAY
switch is turned to the position in
which new data is required and the
KYBD pushbutton is pressed, data
may be displayed on the appropriate left, right, and center display.
To display a number, press the corresponding key or keys (1 through
0). To display a letter, first depress
the key corresponding to the desired letter. Then depress a key in
the left, middle or right column,
corresponding to the position of the
letter on the key. Example: To enter an L, first depress L, then 3, 6,
or 9 in the right column.

FLY-TO-DEST
thumbwheel
switch

Selects the destination for which
XTK/TKE and DIST/BRG/TIME
are displayed when the DISPLAY
switch is turned to either of these
positions which steering information is desired. Destinations are 0
through 9, and H (Home).

ENT key

Enters data set up on keyboard into
memory when pressed.

CLR key

Clears last entered character when
pressed once. When pressed twice,
clears entire display panel under
keyboard control.

Spheroid code of destination set on
DEST DISP thumbwheel.
Magnetic variation (in degrees and
tenths of degrees) of destination set
on DEST DISP thumbwheel.

MEM indicator
lamp

Lights when radar portion of navigation set is in nontrack condition.

MAL indicator
lamp

Lights when navigation set malfunction is detected by built in selftest.

DIM control

Controls light intensity of display
characters.

Left, Right, and
Center display
lamps

Lights to provide data in alphanumeric and numeric characters, as
determined by setting of DISPLAY
switch, MODE switch, and operation of keyboard.

Target storage
indicator

Displays
destination
number
(memory location in which present
position will be stored when TGT
STR pushbutton is pressed.

TGT STR
pushbutton

Stores present position data when
pressed.

FUNCTION

3-39

TM 1-1520-237-10

3.17.3 Modes of Operation. The three basic modes of
operation are: Navigate, test, and backup.
3.17.3.1 Test Mode. The TEST mode contains two
functions: LAMP TEST mode, in which all display segments are lit, and TEST mode, in which system operation
is verified. In the LAMP TEST mode, system operation is
identical to that of the navigate mode except that all lamp
segments and the MEM and MAL indicator lamps are
lighted to verify their operation (Figure 3-18). In TEST
mode, the system antenna no longer transmits or receives
electromagnetic energy; instead, self-generated test signals
are inserted into the electronics to verify operation. System
operation automatically reverts into the backup mode during test mode. Self-test of the Doppler set is done using
built-in-test equipment (BITE), and all units connected and
energized for normal operation. Self-test isolates failures to
one of the three units. The computer-display unit (except
for the keyboard and display) is on a continuous basis, and
any failure is displayed by turn-on of the MAL indicator
lamp on the computer-display unit. The signal data converter and receiver-transmitter-antenna are tested by turning the MODE switch to TEST. Failure of those components is displayed on the computer-display unit by turn-on
of the MAL indicator lamp. Identification of the failed unit
is indicated by a code on the display panel of the computerdisplay unit. Continuous monitoring of the signal data converter and receiver-transmitter-antenna is provided by the
MEM indicator lamp. The MEM indicator lamp will light
in normal operation when flying over smooth water. However, if the lamp remains on for over 10 minutes, over land
or rough water, there is a malfunction in the Doppler set.
Then the operator should turn the MODE switch to TEST,
to determine the nature of the malfunction. Keyboard operation is verified by observing the alphanumeric readout as
the keyboard is used.
3.17.3.2 Navigate Mode. In the navigate mode (UTM
or LAT/LONG position of the MODE selector), power is
applied to all system components, and all required outputs
and functions are provided. Changes in present position are
computed and added to initial position to determine the
instantaneous latitude/longitude of the helicopter. Destination and present position coordinates can be entered and
displayed in UTM and latitude/longitude. At the same time,
distance, bearing and time-to-go to any one of ten preset
destinations are computed and displayed as selected by the
FLY-TO DEST thumbwheel.
3.17.3.3 Backup Mode. In this mode, remembered velocity data are used for navigation. The operator can insert
ground speed and track angle with the keyboard and the
display in GS-TK position. This remembered velocity data
can be manually updated through use of the keyboard and

3-40

CDU DISPLAY switch in the GS-TK position. When
GS-TK values are inserted under these conditions, navigation continues using only these values.
3.17.4 Operation.
3.17.4.1 Window Display and Keyboard Operation.
In all data displays except UTM coordinates, the two fields
are the left and right display windows. In UTM coordinates
displays, the first field of control is the center window and
the second field is the combination of the left and right
displays. When pressing the KYBD pushbutton, one or
other of the fields described above is under control. If it is
not desired to change the display in the panel section under
control, the pilot can advance to the next field of the display
panel by pressing the KYBD pushbutton again. The last
character entered may be cleared by pressing the CLR key.
That character may be a symbol or an alphanumeric character. However, if the CLR key is pressed twice in succession, all characters in the field under control will be cleared
and that field will still remain under control.
3.17.4.2 Data Entry.
1. To enter a number, press the corresponding key.
To enter a letter, first press the key corresponding to the desired letter. Then press a key in the
left, middle, or right column corresponding to
the position of the letter on the pushbutton.
2. Example: To enter an L, first press L, then either 3, 6, or 9 in the right column. The computer program is designed to reject unacceptable data (for example, a UTM area of WI does
not exist, and will be rejected). If the operator
attempts to insert unacceptable data, the display
will be blank after ENT is pressed.
3.17.4.3 Starting Procedure.
1. MODE selector - LAMP TEST. All lights
should be lit.
a. Left, right, Center and Target storage indicator - Lit (Figure 3-18). All other lights
should be on.
b. Turn DIM control fully clockwise, then
fully counterclockwise, and return to full
clockwise; all segments of the display
should alternately glow brightly, go off,
and then glow brightly.
2. MODE selector - TEST. After about 15 seconds left display should display GO. Ignore the

TM 1-1520-237-10

LEFT
DISPLAY
LAMPS

RIGHT
DISPLAY
LAMPS

TENTHS OF MINUTES
DECIMAL

KYBD

MEM MAL

TGT
STR

TARGET
STORAGE
INDICATOR

CENTER
DISPLAY
LAMPS

AA0525
SA

Figure 3-18. Doppler Lamp Test Mode Display
random display of alpha and numeric characters which occurs during the first 15 seconds.
Also ignore test velocity and angle data displayed after the display has frozen. After about
15 seconds, one of the following five displays
will be observed in the first two character positions in the left display:
NOTE

DISPLAY
LEFT

RIGHT

No display.
Display
blanks
(normal).

REMARKS

If right display is blank,
system
is
operating
satisfactorily.

If the MAL lamp lights during any mode of
operation except LAMP TEST, the
computer-display unit MODE switch should
be turned first to OFF, and then to TEST, to
verify the failure. If the MAL lamp remains
on after recycling to TEST, notify organizational maintenance personnel of the navigation set malfunction.

3-41

TM 1-1520-237-10

DISPLAY
LEFT

RIGHT

DISPLAY

REMARKS

If right display is P, then
pitch or roll data is missing,
or pitch exceeds 90°. In this
case, pitch and roll in the
computer are both set to zero
and navigation continues in a
degraded operation. Problem
may be in the vertical gyroscope or helicopter cabling.

LEFT

RIGHT

C, R, S, or H
followed
by a numeric
code

A failure has occurred and the
BACKUP mode, used for
manual navigation (MN), is
the only means of valid navigation. The operator may use
the computer as a dead reckoning device by entering
ground speed and track data.
The operator should update
present position as soon as
possible, because it is possible significant navigation
errors may have accumulated.

HO10000

No heading information to
signal data converter.

SO5000

No 26 vac to signal data converter.

C, R, S, or H
followed
by a numeric
code

A failure has occurred in the
system and the operator
should not use the system.

NOTE

If the TEST mode display is MN or NG, the
MODE switch should be recycled through
OFF to verify that the failure is not a momentary one. If the TEST mode display is MN,
the data entry may be made in the UTM or
LAT/LONG mode, but any navigation must
be carried on with the system in the
BACKUP mode.
BU

3-42

C, R, S, or H
followed
by a numeric
code

A failure has occurred and the
system has automatically
switched to a BACKUP
mode of operation as follows:
1. The operator has the option
of turning the MODE switch
to BACKUP and entering the
best estimate of ground speed
and track angle. 2. The operator has the option of turning
the MODE switch to
BACKUP and entering his
best estimate of wind speed
and direction and entering his
best estimate of ground speed
and track angle. The operator
should update present position as soon as possible, because it is possible that significant navigation errors may
have accumulated.

REMARKS

The 9V battery has failed. All
stored data must be reentered
after battery replacement.

Blank

C with
random
numbers

Computer display unit failure.

Blank

R with
random
numbers

Receiver-transmitter-antenna
failure.

Blank

S with
random
numbers

Signal data converter failure.

Random
display

Random display

Signal data converter failure.

TM 1-1520-237-10

3.17.4.4 Entering UTM Data. This initial data is inserted before navigating with the Doppler. Refer to paragraph 3.17.4.9.
a. Spheroid of operation, when using UTM coordinates.
b. UTM coordinates of present position - zone, area,
easting (four significant digits) and northing (four significant digits; latitude/longitude coordinates may be used.
c. Variation of present position to the nearest one-tenth
of a degree.
d. Coordinate of desired destination - 0 through 5 and
H; (6 through 9 are normally used for target store locations;
but may also be used for destinations). It is not necessary to
enter all destinations in the same coordinate system.

6. ENT pushbutton - Press if no variation data is
to be entered.
7. KYBD pushbutton - Press, if variation data is
to be entered, and note right display blanks. (If
no variation data is to be entered, ENT key Press.)
8. Variation data - Enter. (Example: E001.2, press
keyboard keys 2 (right window blanks), 2, 0, 0,
1 and 2. Press ENT key, the entire display will
blank and TGT STR number will reappear,
display should indicate INø E 001.2.)
3.17.4.6 Entering Present Position or Destination
In UTM.
1. MODE selector - UTM.

NOTE
2. DISPLAY selector - DEST-TGT.
It is not necessary to enter destinations unless steering information is required, unless
it is desired to update present position by
overflying a destination, or unless a present
position variation computation is desired
(paragraph 3.17.3.3). If a present position
variation running update is desired, destination variation must be entered. The operator
may enter one or more destination variations
to effect the variation update; it is not necessary for all destinations to have associated
variations entered.
3.17.4.5 Entering Spheroid and/or Variation.
1. MODE selector - UTM, LAT/LONG or
BACKUP.
2. DISPLAY selector - SPH-VAR.
3. DEST DISP thumbwheel - P, numeral, or H as
desired.
4. KYBD pushbutton - Press. Observe display
freezes and TGT STR indicator blanks. Press
KYBD pushbutton again and observe left display blinks. If no spheroid data is to be entered,
KYBD pushbutton -Press again, go to step 7.
5. Spheroid data - Entry. (Example: INø). Press
keys 3 (left window blanks), 3, 5, 5 and 0. Left
display should indicate INø. Refer to Figure
3-22 for codes.

3. DEST DISP thumbwheel - P, numerical, or H
as desired.
4. Present position and destination - Enter. (Example: Entry of zone 31T, area CF, easting
0958 and northing 3849.)
a. KYBD pushbutton - Press. Observe that
display freeze and TGT STR indicator
blanks.
b. KYBD button - Press. Observe that center
display blanks.
c. Key 3, 1, 7, and 8 - Press.
d. KYBD button - Press. Observe left and
right displays blank.
e. Key 1, 3, 2, 3, 0, 9, 5, 8, 3, 8, 4, 9 - Press.
f. ENT pushbutton - Press. Left, right, and
center displays will momentarily blank and
TGT STR number will appear. Displays
should indicate 31T CF 09583849.
3.17.4.7 Entering Present Position or Destination
Variation In LAT/LONG. The variation of a destination
must be entered after the associated destination coordinates
are entered (since each time a destination is entered its
associated variation is deleted). The order of entry for
present position is irrelevant.

Change 1

3-43

TM 1-1520-237-10

NOTE
If operation is to occur in a region with relatively constant variation, the operator enters
variation only for present position, and the
computer will use this value throughout the
flight.
1. MODE selector - LAT/LONG.
2. DISPLAY selector - DEST-TGT.
3. DEST DISP thumbwheel - P, numerical or H
as desired.
4. Present position or destination - Enter. (Example: Entry of N41° 10.1 minutes and E035°
50.2 minutes.) Press KYBD pushbutton. Observe that display freezes and TGT STR indicator blanks. Press KYBD pushbutton again
and observe left display blanks. Press keys 5, 5,
4, 1, 1, 0 and 1. Press KYBD pushbutton (right
display should clear), and keys 2, 2, 0, 3, 5, 5, 0
and 2.

2. DISPLAY selector - DEST-TGT.
3. DEST DISP thumbwheel - P. Do not press
ENT key now.
4. ENT pushbutton - Press as helicopter is sitting
over or overflies initial fix position.
5. FLY-TO DEST thumbwheel - Desired destination location.
3.17.4.10 Update of Present Position From Stored
Destination. The helicopter is flying to a destination set
by the FLY-TO DEST thumbwheel. When the helicopter
is over the destination, the computer updates the present
position when the KYBD pushbutton is pressed, by using
stored destination coordinates for the destination number
shown in FLY-TO DEST window, and adding to them the
distance traveled between the time the KYBD pushbutton
was pressed and the ENT key was pressed.
1. DISPLAY selector - DIST/BRG-TIME.
2. KYBD pushbutton - Press, when helicopter is
over the destination. Display freezes.

5. ENT pushbutton - Press. Entire display will
blank and TGT STR number will reappear.
Display should indicate N 41° 10.1 E0 35° 50.2.
3.17.4.8 Ground Speed and Track.
1. MODE selector - BACK UP.

NOTE
If a present position update is not desired, as
indicated by an appropriately small value of
distance to go on overflying the destination,
set the DISPLAY selector to some other position, this aborts the update mode.

2. DISPLAY selector - GS-TK.
3. ENT key - Press.
3. Ground speed and track - Enter. (Example: Enter 131 km/h and 024°. Press KYBD pushbutton, observe that left display freezes and TGT
STR indicator blanks. Press KYBD pushbutton
and observe that left display blanks. Press keys
1, 3, and 1. Left display indicates 131. Press
KYBD pushbutton, control shifts to right display, and right display blanks. Press keys 0, 2
and 4.
4. ENT pushbutton - Press. The entire display will
blank, and TGT STR number will reappear.
Display should indicate 131 024°.
3.17.4.9 Initial Data Entry. Initial data entry of variation in coordinates is normally done prior to takeoff. To
make the initial data entry, do the following:
1. Present positon variation - Enter (paragraph
3.17.4.5).

3-44

3.17.4.11 Update of Present Position from Landmark. There are two methods for updating present position from a landmark. Method 1 is useful if the landmark
comes up unexpectedly and the operator needs time to determine the coordinates. Method 2 is used when a landmark
update is anticipated.
a. Method 1.
(1) DISPLAY selector - PP.
(2) KYBD pushbutton - Press as landmark is
overflown. Present position display will
freeze.
(3) Compare landmark coordinates with those
on display.
(4) Landmark coordinates - Enter. If difference
warrants an update.

TM 1-1520-237-10

(5) ENT key - Press if update is required.
(6) DISPLAY selector - Set to some other position to abort update.
b. Method 2.
(1) DISPLAY selector - DEST/TGT.

(1) MODE selector - UTM or LAT/LONG,
depending on coordinate format desired.
(2) DISPLAY selector - DEST-TGT.
(3) DEST DISP thumbwheel - P.
(4) KYBD pushbutton - Press when over flying potential target. Display should freeze.

(2) DEST DISP thumbwheel - P. Present postion coordinate should be displayed.

NOTE

(3) KYBD pushbutton - Press, observe that
display freezes.

Do not press ENT key while DEST DISP
thumbwheel is at P.

(4) Landmark coordinates - Manually enter via
keyboard.
(5) ENT key - Press when overflying landmark.
(6) DISPLAY selector - Set to some other position to abort update.
3.17.4.12 Left-Right Steering Signals. Flying shortest distance to destination from present position.
1. DISPLAY selector - XTK-TKE.
2. MODE SEL - DPLR.
3. Fly helicopter in direction of lateral deviation
pointer on vertical situation indicator to center
the pointer, or course deviation bar on HSI.
3.17.4.13 Target Store (TGT STR) Operation. Two
methods may be used for target store operation. Method 1
is normally used when time is not available for preplanning
a target store operation. Method 2 is used when time is
available and it is desired to store a target in a specific
DEST DISP position.

(5) If it is desired to store the target, turn
DEST DISP thumbwheel to destination location desired and press ENT key.
(6) If it is not desired to store the target, place
DISPLAY selector momentarily to another
position.
3.17.4.14 Transferring Stored Target Coordinates
From One Location to Another. The following procedure allows the operator to transfer stored target coordinates from one thumbwheel location to another. For example, it is assumed that the pilot wants to put the
coordinates of stored target 7 into location of destination 2.
NOTE
Throughout this procedure, range, time-togo, bearing and left/right steering data are
computed and displayed for the destination
selected via the FLY-TO DEST thumbwheel.
1. DISPLAY selector - DEST-TGT.
2. DEST DISP thumbwheel - 7.
3. KYBD pushbutton - Press.

a. Method 1.
4. DEST DISP thumbwheel - 2.
(1) TGT STR pushbutton - Press when flying
over target.
(2) Present position is automatically stored and
the destination location is that which was
displayed in the target store indicator (position 6, 7, 8, or 9) immediately before
pressing the TGT STR pushbutton.
b. Method 2.

5. ENT key - Press.
3.17.4.15 Transferring Variation From One Location to Another. The procedure to transfer variation data
to the same location where the associated stored target coordinates has been transferred is the same as in paragraph
3.17.4.14. Transferring Stored Target Coordinates From
One Location To Another, except that the DISPLAY selector is placed at SPH-VAR.

3-45

TM 1-1520-237-10

3.17.4.16 Dead Reckoning Navigation. As an alternate BACKUP mode, dead reckoning navigation can be
done using ground speed and track angle estimates provided by the operator.
1. MODE selector - BACKUP.

able in both Military Grid Reference System (MGRS) and
Latitude and Longitude (LAT/LONG) coordinates. Navigation and steering is performed using LAT/LONG coordinates and a bilateral MGRS-LAT/LONG conversion routine is provided for MGRS operation. Up to 100
destinations may be entered in either format and not necessarily the same format.

2. DISPLAY selector - GS-TK.
3. Best estimate of ground speed and track angle Enter via keyboard.
4. Set MODE selector to any other position to
abort procedure.
3.17.4.17 Operation During and After Power Interruption. During a dc power interruption inflight, or when
all helicopter power is removed, the random access memory
(RAM) (stored destination and present position) data is retained by power from an 8.4 volt dc dry cell battery. This
makes it unnecessary to reenter any navigational data when
power returns or before each flight. If the battery does not
retain the stored destination data during power interruption,
the display will indicate on EN when power returns. This
indicates to the pilot that previously stored data has been
lost, and that present position, spheroid/variation, and destinations must be entered. The computer, upon return of
power, resets present position variation to E000.0°, destination and associated variations to a non-entered state, remembers wind to zero and spheroid to CL6. The following
data must be entered following battery failure:

3.17A.1 Antenna. The GPS antenna is located on the
top aft section of the helicopter. The Doppler antenna is
located below the copilot’s seat (Figure 3-1).
3.17A.2 Controls, Displays, and Function. The control and displays for the AN/ASN-128B are on the front
panel (Figure 3-18.1). The function of each control is as
follows:
CONTROL/
INDICATOR

NOTE

The MODE switch is locked in the
OFF position and must be pulled
out and turned to get into or out of
the OFF position.
MODE selector

Selects mode of operation.

OFF

In this position the navigation set is
inoperable: non-volatile RAM
retains stored waypoint data.

1. Enter spheroid.

LAMP TEST

Checks operation of all lamps.

2. Enter present position variation.

TEST

Initiates built-in-self test exercise
for the Doppler and GPS functions
of the navigation set.

MGRS

Selects MGRS navigational mode
of operation.

LAT/LONG

Selects
latitude/longitude
navigational mode of operation.

GPS LDG

Places navigation set in GPS
landing mode of operation. This
mode provides real time, tactical
precision
landing
guidance
information to the HSI and VSI
indicators.

3. Enter present position.
4. Enter each destination and its associated variation.
3.17.5 Stopping Procedure. MODE selector - OFF.
3.17A DOPPLER/GPS NAVIGATION SET (DGNS)
AN/ASN-128B. UH
The AN/ASN-128B DGNS is an AN/ASN-128 LDNS
with an embedded GPS receiver. The AN/ASN-128B in
conjunction with the aircraft’s heading, vertical references,
and position and velocity updates from its internal GPS,
provides accurate aircraft velocity, position and steering information from ground level to 10,000 feet. The system
provides worldwide navigation, with position readout avail-

3-46

FUNCTION

Change 4

DISPLAY
selector

Selects navigation data for display.

TM 1-1520-237-10

FLY TO

BRT

DIM

G
P
S
/
D
P
L
R

EPE

TGT
STR

1 7 : BANDO 0 3 0MG 9 1
GP S : M NA V : C
GS : 1 1 7 KM / HR
TK : 0 2 5 "

PP
GS/TK
NAV M

DIST / BRG
TIME
WP
TGT

XTK/TKC
KEY
WIND−UTC
DATA

MAL

LTR

LEFT

MID

RIGHT

ABC
1

DEF
2

GHI
3

TGT
STR

JKL
4

MNO
5

PQR
6

INC
(+)

STU
7

VWX
8

YZ*
9

DEC
(−)

CLR

#
0

(PAGE)

KYBD

DATUM
ROUTE

DISPLAY
TEST

N
A
V

SYS
STAT

MGRS
LAT /
LONG

LAMP
TEST

GPS
LDG

OFF

MODE

ENT

AA9998A
SA

Figure 3-18.1. Doppler/GPS Navigation Set
AN/ASN-128B

FUNCTION

CONTROL/
INDICATOR

FUNCTION

WIND-UTC
DATA 9Doppler
only9

Used for wind speed and direction
(with TAS sensor installed), UTC
time, sea current, surface wind,
GPS status and data load functions.

DIST BRG
TIME

XTK/TKE
KEY

Displays steering (cross track distance and track angle error) information and GPS variable key status. Selection of fly to destination
by direct entry of two digit destination number.

Displays distance, bearing and time
information to the destination or
course selected. Selection of fly to
destination can be accomplished by
direct entry of two digit destination
number.

WP TGT

Accesses waypoint or target data
(landing data, variation, motion).
Selection of destination for display/
entry by direct entry of two digit
destination number.

DATUM
ROUTE

Accesses datum and steering/route
functions.

CONTROL/
INDICATOR

GS/TK NAV M

Displays ground speed, track angle
and selection of GPS and navigation mode.
Displays present position, altitude
and magnetic variation.

MAL indicator
lamp

Lights when a malfunction is detected by the built-in-test circuitry.
In the event of an intermittent malfunction, the system may operate
correctly but must be cycled to the
OFF position then to on, to extinguish the MAL light.

Change 2

3-46.1

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

CONTROL/
INDICATOR

FUNCTION

BRT and DIM
keys

Used to brighten or dim the light
intensity of the LCD display.

INC and DEC
keys

Four line alphanumeric display

Displays alphanumeric characters,
as determined by the setting of the
DISPLAY selector, the MODE selector and operation keyboard. The
keys activate function upon pressing the key.

TGT STR key

Stores present position data in the
indicated target store/memory location (90-99) when pressed.

Used to increment or decrement the
displayed waypoint/target number
when the DISPLAY selector is set
to WP/TGT. To access P, press the
LTR LEFT key followed by key 6;
display waypoint 99 then press the
INC key; or display waypoint 00
then press the DEC key. Also used
to increment or decrement the fly-to
destination number when the DISPLAY selector is set to DIST/
BRG/TIME or XTK/TKE/KEY.

KYBD key

Used in conjunction with the keyboard to allow data display and entry into the computer.

ENTkey
(PAGE)

Keyboard and
LTR keys

Used to set up data for entry into
memory. When DISPLAY selector
is set to a position in which new
data is required and KYBD key is
pressed, data may be displayed on
the appropriate input field of display. To display a number, press the
corresponding key or keys (0-9). To
display a letter, first press the LTR
key corresponding to the position of
the desired letter on a key. Then
press the key which contains the
desired letter. Example: To enter an
L, first press the LTR RIGHT key,
then press key 4.

Enters data into memory (as set up
on keyboard and displayed). This
key is also used for paging of displays. The bottom right corner of
the display indicates 9more9 when
additional pages are available, and
9end9 when no additional pages are
available. Pressing this key when
9end9 is displayed will return the
display to the first page.

CLR key

Clears last entered character when
pressed once. When pressed twice,
clears entire input field of display
keyboard control.

F1 key

Reserved for future growth.

3-46.2

Change 2

TM 1-1520-237-10

3.17A.3 Modes of Operation. Control of the Doppler/
GPS, including selection of modes and displays, and entry
and readout of data is performed via the Computer Display
Unit (CDU) front panel. The system has four basic modes
of operation: OFF, navigate, TEST and GPS LDG. In the
navigate mode three submodes may be selected manually
or automatically. These are combined mode (default or primary mode of operation), GPS only mode, or Doppler only
mode.
3.17A.3.1 OFF Mode. In the OFF mode the system is
inoperable. However, the edge lighting is lighted by an external aircraft power source and is independent of the
Doppler/GPS MODE selector setting. Edge lighting may
not be available if the helicopter is modified with the night
vision MWO.
3.17A.3.2 Navigate Mode. In the navigate mode
(MGRS or LAT/LONG) position of the CDU MODE selector) power is applied to all system components, and all
required outputs and functions are provided. The Doppler
radar velocity sensor (DRVS) measures aircraft velocity,
and converts analog heading, pitch and roll into digital
form. This data and embedded GPS receiver (EGR) velocity and position data are then sent to the CDU for processing. Barometric altitude is used for aiding the GPS when
only three satellites are available. Four satellites are required if the barometric altitude sensor is not available.
Present position is computed by using one of three navigation submodes which can be selected manually or automatically. These submodes are as follows:
3.17A.3.2.1 Combined Mode (Default or Primary
Mode of Operation). Doppler and GPS position and velocity data are combined to provide navigation. This mode
is used when a minimum of three (with barometric sensor)
or four satellites are available, GPS Estimated Position Error (EPE) is less than approximately 150 meters, and the
Doppler is not in memory. If GPS becomes invalid (e.g.
due to increased EPE), the system will automatically switch
to Doppler mode until a valid GPS status is received. The
GPS POS ALERT advisory light will illuminate when this
happens. If the Doppler becomes invalid (e.g. flight over
glassy smooth water), the system will automatically switch
to GPS mode if GPS is valid or an alternate Doppler mode
if the GPS is not valid.
3.17A.3.2.2 GPS Mode. GPS positions and velocities
are used for navigation by the Doppler navigation processor
in the CDU. If GPS mode is selected and the GPS becomes
invalid (paragraph 3.17A.3.2.1), the system will not navigate. The GPS POS ALERT advisory light indicates that
GPS signals are not reliable.

3.17A.3.2.3 Doppler Mode. Doppler position and velocity data are used for navigation. If Doppler mode is selected and the Doppler becomes invalid (paragraph
3.17A.3.2.1), the system will automatically switch to True
Air Speed (TAS) mode (using remembered wind) if a TAS
sensor is available, or remembered velocity if a TAS sensor
is not available. If Doppler mode is manually selected at
the start of the flight an initial present position must be
obtained and entered prior to flight. Navigation is performed in latitude/longitude for computational convenience
only. At the same time, distance, bearing and time-to-go to
any one of 100 preset destinations are computed (as selected by FLY-TO-DEST).
3.17A.3.3 Test Mode. The TEST mode contains two
functions: LAMP TEST mode, in which all display segments are lit, and TEST mode, in which system operation
is verified. In the LAMP TEST mode, system operation is
identical to that of the navigate mode except that all lamp
segments and the MEM and MAL indicator lamps are
lighted to verify their operation. In TEST, the RTA no
longer transmits or receives electromagnetic energy; instead, self-generated test signals are inserted into the electronics to verify operation of the DRVS. At this time a self
test is performed by the GPS and navigation computations
continue using remembered velocity. In the TEST mode,
Doppler test results are displayed on the CDU front panel
for the first 15 seconds (approximate). At the end of this
period either GO is displayed if there is no malfunction in
the navigation set, or a failure code is displayed if a malfunction has occurred. A rotating bar on the display indicates that the GPS has not completed self test. If the navigation set is maintained in the TEST mode, no navigation
data can be displayed on the CDU front panel. If a Doppler
malfunction is detected, the MAL indicator lamp lights and
DF is displayed. At the completion of GPS self test (up to
two minutes), the rotating bar is replaced with a complete
test result code. The failed unit and the failed circuit card
are also indicated by a code on the CDU display. The CDU
is continuously monitored for failures, using its own computer as built-in-test-equipment (BITE). Any BITE malfunction causes the MAL indicator lamp on the CDU to
light. If the MODE selector on the CDU is set to TEST,
identification of the failed LRU is indicated by a code on
the display panel. Aircraft heading, pitch and roll are also
displayed in the mode by pressing the ENT key after Doppler test is completed. GPS test status is displayed if the
ENT key is pressed a second time. Malfunction codes are
automatically latched and can only be cleared by recycling
the CDU power via the CDU mode switch (OFF-ON).
3.17A.3.4 GPS Landing Mode. In the GPS LDG
mode, the Doppler navigation system provides information

Change 8

3-46.3

TM 1-1520-237-10

to the HSI and VSI indicators for real time landing guidance to a touch down point previously entered in any of the
100 fly-to destinations. The landing approach is determined
by present position and the entered touch down altitude,
glideslope and inbound approach course.
3.17A.4 CDU Operation. Various required operating
data, such as initial present position (if GPS is not valid or
Doppler mode is selected), destination coordinates with or
without GPS landing data, and magnetic variation can at
any time be entered into the CDU via its keyboard, or the
data loader (Figure 3-11) via the preprogrammed data
loader cartridge. In most cases, these data will be entered
before the aircraft takes off. The GPS provides present position to the Doppler/GPS. If GPS is not available or Doppler is selected present position can be initialized as follows:
1. The MODE selector should be set to MGRS or
LAT/LONG, the WP/TGT display position of
the DISPLAY selector is selected, the destination number is set to P (default waypoint) and
KYBD key is pressed. The coordinates of the
initial position is overflown, the ENT key is
pressed. The computer then determines changes
from the initial position continuously, and the
coordinates of the current present position can
be read either by remaining in this configuration or by setting the DISPLAY selector to PP
(present position) and the MODE selector to
MGRS or LAT/LONG.
2. To update present position over a stored destination, KYBD key is pressed when the aircraft
overflies this destination. If an update is desired, the ENT key is pressed and the update is
completed. The DISPLAY selector is in the
DIST/BRG/TIME position and the FLY-TODEST is set to this destination during this process. The distance-to-go, displayed while over
the stored destination, is the position error of
the system at that moment.
3. To update present position over a fixed point
not previously stored in the computer, the DISPLAY selector is placed to PP and KYBD key
is pressed as the fix point is overflown. This
freezes the display while allowing computation
of changes in present position to continue
within the computer. If an update is required
the coordinates of the fix point are entered via
the keyboard, and ENT key is pressed. The position change which occurred since over-flying
the fix point is automatically added to the fix

3-46.4

Change 10

point coordinates to complete the position update.
4. Magnetic variation can be entered for each destination, and the system will compute present
position magnetic variation. If operation is to
occur in a region with relatively constant magnetic variation, the operator enters magnetic
variation only for present position and the computer will use this value throughout the flight. If
MGRS data are to be entered or displayed, the
MGRS datum of operation is also entered.
3.17A.5 Target-of-Opportunity. Target-of-opportunity
data can be stored by pressing TGT STR (target store) key
when the target is overflown. This operation stores the coordinates of the target in one of ten destination locations in
the computer; locations 90-99 sequentially incrementing
each time the TGT STR key is pressed. The location is
displayed in the appropriate display field. The computer
can keep track of individual target positions which may
include speeds and directions input by the operator.
3.17A.6 Self Test. Self test of the AN/ASN-128B is accomplished using BITE with the RTA, SDC, and CDU
units connected and energized for normal operation. Self
test enables the unit to isolate failures to one of the four
main functions (RTA, SDC, CDU or EGR) or to one of
the circuit cards in the SDC or CDU. Self test is accomplished as follows:
1. The CDU (except for the keyboard and display)
is checked on a continuous basis, and any failure is displayed by the illumination of the MAL
indicator lamp on the CDU. If the MODE selector on the CDU is set to the TEST position,
identification of the failed circuit card in the
CDU is indicated by a code on the display
panel.
2. The DRVS and EGR are tested by setting the
MODE selector on the CDU to the TEST position. Failure of the DRVS or EGR are displayed on the CDU by illumination of the MAL
indicator lamp, and identification of the failed
unit or circuit card is indicated by a code on the
display panel of the CDU.
3. Continuous monitoring of the signal data converter and receiver transmitter antenna is provided by the system status indication. The system will not use Doppler velocities in normal
operation when flying over glassy smooth water. However, if the system continues to not use

TM 1-1520-237-10

Doppler (e.g. using GPS only when combined
has been selected) for excessive periods of time
(e.g. more than 10 minutes) over land or rough
water, then a malfunction may exist in the navigation set and the operator should set the
MODE selector to TEST to determine the nature of the failure.
4. The display portion of the CDU is tested by
illuminating all the lamp segments in each alphanumeric character in the LAMP TEST
mode.
5. Keyboard operation is verified by observing the
alphanumeric characters as the keyboard is exercised.
3.17A.7 Route Sequencing Modes. The system has
the ability to fly a preprogrammed sequence of waypoints.
This sequence can be either consecutively numbered in
which case a start and end waypoint are entered or random
numbered, in which case all waypoints are put in a list and
the start and end waypoints are entered. Both sequence
modes can be flown in the order they are in the list or in the
reverse order. Directions will be displayed to the waypoint
next on the list until approximately 10 seconds before overflying the waypoint at which time the display will advance
to the next waypoint and the new waypoint number will
blink for ten seconds. One consecutive and one random
sequence may be stored in the system.
3.17A.8 To-To Route Mode. The system has the ability
to provide steering information onto a course defined by the
start and end waypoints. Only the second waypoint will be
overflown. The distance displayed is the distance to the
course when outside two nautical miles of the course and
the distance to the second waypoint when inside two nautical miles of the course.
3.17A.9 General Operating Procedures for Entering Data. The panel display consists of four line LED
readout. The top line of the display is reserved for the display of Fly-To destination number and destination name/
International Civil Aeronautic Organization (ICAO) identifier, EPE in meters, mode of GPS and mode of AN/ASN128B operation and target store number. The remaining
lines will display data in accordance with the DISPLAY
and MODE selectors. When pressing the KYBD key for
the first time in an entry procedure, the display freezes,
kybd is displayed in the bottom right corner indicating the
display is in the keyboard mode and the input field under
keyboard control blinks. If it is not desired to change the
display field under control, the pilot can advance to the next
field of the display by pressing the KYBD key again. Press-

ing the ENT key (whether or not new data has been entered) causes the display to blank momentarily and return
with the latest computed data. To abort a keyboard operation, move the MODE or DISPLAY selector to another
position.
a. Data Entry. To display a letter, first press the LTR
key corresponding to the position of the desired letter on a
key. Then press the key which contains the desired letter.
For example, to enter an L, first press the LTR RIGHT
key, then press key 4.
b. Keyboard Correction Capability. The last character
entered may be cleared by pressing the CLR key. If the
CLR key is pressed twice in succession, the field is cleared
but remains under control (indicated by blinking) and the
last valid data entered is displayed.
c. Destination Variation Constraint. The magnetic variation associated with a destination must be entered after the
coordinates for that destination are entered. The order of
entry for present position is irrelevant.
d. Impossibility of Entering Unacceptable Data. In most
cases the computer program will reject unacceptable data
(for example, a MGRS area of W1 does not exist and will
be rejected). If the operator attempts to insert unacceptable
data, the unacceptable data will be displayed on the panel
and then the selected field will blink after ENT key is
pressed displaying the last valid data.
NOTE
The computer cannot prevent insertion of erroneous data resulting, for example, from
human or map errors.
e. Procedure for Displaying Wind Speed and Direction
(TAS Sensor Required).
NOTE
In MGRS mode, wind speed is displayed in
km/hr; in LAT/LONG mode, wind speed is
displayed in knots. Wind direction is defined
as the direction from which the wind originates.
(1) Set MODE selector to LAT/LONG (MGRS
may also be used).
(2) Set DISPLAY selector to WIND-UTC (coordinated universal time)/DATA and observe
display.

Change 10

3-46.5

TM 1-1520-237-10

(2) Set the MODE selector to LAT/LONG
(MGRS may also be used).

(3) The display indicates:
SP:XXXKn

(3) The display indicates GPS daily key status,
time remaining on the currently entered keys
and how many satellites are currently being
used by the GPS.

DIR:XXX°
f. Procedure for displaying/entering UTC and displaying GPS status.
(1) Set MODE selector to LAT/LONG (MGRS
may also be used).

KEY

STATUS

TIME

REMARKS

Days or hours
still available
on key

GPS daily key
in use and
verified

No GPS daily
key available

GPS daily key
available but
not verified

(2) Set DISPLAY selector to WIND-UTC/DATA
and observe the wind speed/direction display.
(3) Press ENT key. Observe that the CDU display
indicates year ** (default year is 93), day 317
and indicates hours, minutes, and seconds of
UTC time: 09 Hours, 25 Minutes, 10 Seconds.
(4) To enter year, day and time press the KYBD
key to select the field for input shown as a
blinking field, enter the desired data and press
the ENT key.

3.17A.10. Preflight Procedures.
a. Data required prior to DGNS turn-on.

(5) To display GPS status press the ENT key to
display selection menu.
1>SEA CURRENT
2>SURFACE WIND
3>GPS STATUS
4>DATA LOAD end
(6) To select the GPS STATUS page press key 3.
(7) Observe the CDU display. The display indicated the GPS test mode status as of one of the
following:
GPS TEST: IN PROCESS
GPS TEST: NOT RUN
GPS TEST: PASSED
GPS TEST: FAILED
g. Procedure for displaying GPS key and GPS satellite
status.
(1) Set the DISPLAY selector to XTK/TKE/KEY.

3-46.6

Change 2

(1) The following initial data must be entered
by the pilot after system turn-on and initialization, unless previously entered data
is satisfactory:
(2) Datums of operation, when using MGRS
coordinates. This data may be part of the
data load if preprogrammed.
(3) In combined or GPS mode the GPS provides preset position. If the Doppler only
mode is selected MGRS coordinates of
present position - zone area, easting and
northing; latitude/longitude coordinates
may also be used to input present position.
This data may be part of the data load if
preprogrammed. Variation of present position to the nearest one-tenth of a degree.
(4) Coordinates of desired destinations 00-99.
It is not necessary to enter all destinations
in the same coordinate system. This data
may be part of the data load if preprogrammed. Destination locations of 70
through 89 are only programmable through
the data loader.
(5) Variation of destinations to the nearest
one-tenth of a degree.

TM 1-1520-237-10

(6) Crypto-key variables necessary to enable
the GPS receiver to operate in Y code are
entered via remote fill data only and not
via the CDU keyboard.
NOTE
Destinations are entered manually when
steering information is required to a destination that was not in the set of data loaded via
the data loader, or it is desired to update
present position by overflying a destination,
or a present position variation computation
is desired. (See CDU operation). If a present
position variation update is desired, destination variation must be entered. The operator
may enter one or more destination variations; it is not necessary for all destinations
to have associated variations entered and
also not necessary to enter all destinations in
any case, but variations must be entered after destination coordinates are entered.
(7) The Doppler outputs true heading and accepts magnetic heading from gyromagnetic
heading reference. If accurate magnetic
variations are not applied, then navigation
accuracy will be affected.
b. System Initialization.
(1) Enter GPS mode 9M 9.

tered, or collection of almanac data when set
has no previous almanac data. During this
time the GPS operation mode must be M
and uninterrupted. After this time the GPS
operating mode may be switched to Y. Observe the GPS key status and number of satellite vehicles (SVs) tracked after switching
to Y mode. If the SV number goes to zero,
repeat this procedure. The key status shall
switch from DK IN to DK OK sometime
during the 12 minutes.
(5) Check datum of operation, if MGRS is being used.
(6) Check destinations in MGRS or LAT/
LONG coordinates as desired.
(7) Check associated destination variations as
desired. Remove all incorrect variations by
setting DISPLAY selector to WP/TGT,
setting the destination number to appropriate destination, and pressing the KYBD
key and ENT key in that order. Variations
of at least two destinations must be entered
for automatic variation update computation
to be performed. For accurate navigation it
is advised to enter variations after each
destination unless the variations are the
same.
(8) Select DGNS operating mode.

NOTE
NOTE
Select GPS mode 9M 9 during initialization.
If 9Y 9 mode is selected before crypto-key
variables are loaded the system will lock-up.
System must be turned off, then back on.

The set will automatically select combined
mode (default or primary operating mode) as
this allows the system to select the best possible navigation method available.

(2) Perform self test.
(3) Perform download of data loader cartridge
if necessary, or manually enter datum, destinations, magnetic variations, and present
position.
(4) Load crypto-key variables (Figure 3-11)
(unless previously loaded and still valid)
necessary for operation of the GPS in Y
mode.

(9) Set the FLY-TO-DEST to the desired destination location.
c. Procedure for downloading data from dataloader cartridge.
(1) Set the CDU MODE selector to OFF.
(2) Insert the preprogrammed data loader cartridge.

NOTE
It is necessary to wait at least 12 minutes for
key validation when new keys have been en-

(3) Set the CDU MODE selector to MGRS
(LAT/LONG may be used). Enter desired
GPS code (M or Y) mode of operation.

Change 10

3-46.7

TM 1-1520-237-10

(4) Set the DISPLAY selector to WINDUTC/DATA.
(5) To display the select menu press the ENT
key twice.

(10) Set the CDU MODE selector to OFF, remove the data loader cartridge if desired,
and then set the CDU MODE selector to
the desired setting.
d. Self-Test.

1>SEA CURRENT
(1) Set the MODE selector to LAMP TEST.
Enter GPS mode 9M 9 or 9Y 9. Verify the
following:

2>SURFACE WIND
3>GPS STATUS

(a) All edge lighting is illuminated.
4>DATA LOAD end

(b) The MAL lamp is illuminated.

(6) To select the DATA LOADER page press
key 4.
DATA LOADER
ENTER DATA: N - Y
(7) To begin the download press the KYBD
and enter Y (yes).

(c) All keyboard keys are lit.
(2) Set the MODE selector to TEST. After
Doppler and/or GPS self tests have completed (approximately 15 seconds for Doppler, up to 2 minutes for GPS), one of the
following displays will be observed in the
left and right displays:
NOTE

(8) Observe the CDU display. The CDU shall
display DOWNLOAD WAYPTS IN
PROCESS. If a transmission error occurs
the CDU display shall change to ERRORRETRYING.
(9) When the transmission is complete the
CDU shall display DOWNLOAD WAYPTS COMPLETE. If this display is not
obtained within one minute of beginning
the download check the data programming
and connections.

LEFT DISPLAY

RIGHT DISPLAY

(3) Press the BRT pushbutton at least 10
times, then press the DIM pushbutton at
least 10 times, then press the BRT pushbutton at least 10 times. LED display shall
alternately glow bright, extinguish, and
glow bright.

REMARKS
Doppler has completed BIT and is operating
satisfactorily, GPS is still performing BIT
(GPS has a two minute BIT cycle maximum). Note that a rotating bar in the display
indicates that the GPS is still performing self
test.

3-46.8

In the event the TEST mode display is not
GO ALL the system should be recycled
through OFF to verify the failure is to a momentary one.

ALL

Change 10

The entire system has completed BIT and is
operating satisfactorily.

TM 1-1520-237-10

(Cont)
LEFT DISPLAY

RIGHT DISPLAY

Pitch or Roll data is missing or exceeds 90°.
In this case, pitch and roll in the computer
are both set to zero and navigation in the
Doppler mode continues with degraded
operation. Problem may be in the vertical
gyro or aircraft cabling.

C, R, S, or H followed by a numeric code

A failure has occurred in the computer
display unit or the signal data converter
power supply. The operator should not use
the system.

GPS failure code

GPS has failed but operator can use Doppler
to perform all navigation.

Doppler failure code

Doppler has failed. GPS is still performing
self test.

Doppler failure code

Doppler has failed but operator can use GPS
to perform all navigation.

Doppler failure code

SDC battery is discharged. Items stored in
memory have been deleted.

e. Procedure for displaying or selecting GPS M
or Y operating mode, Doppler, GPS or combined operation, and displaying groundspeed
and track.

REMARKS

(d) DGNS mode of operation:
C for combined Doppler and GPS.
D for Doppler only.

(1) Set MODE selector to MGRS position
(LAT/LONG or GPS LDG position may
also be used).

G for GPS only.
R for remembered velocities.

(2) Set DISPLAY selector to GS/TK/NAV
M.
(3) The display indicates the current GPS and
navigation mode on the top line:
(a) Selected fly to waypoint.
(b) EPE (GPS estimated position error in
meters). An asterisk (*) in the character position of the EPE display indicates an EPE of greater than 999 or
data unavailable.
(c) GPS mode of operation:
M for mixed C/A and P/Y code GPS reception.
Y for only Y code GPS reception.

* for no navigation.
(e) Target destination where the present
position will be stored next time TGT/
STR is pressed.
NOTE
In MGRS mode, ground speed is displayed
in km/hr; in LAT/LONG mode, ground
speed is displayed in knots.
Only mode C, G, and D may be selected as
the primary navigation mode. Modes R and
* are automatic fall back modes used when
both the Doppler and GPS are unavailable
(4) Selection of GPS mode of operation: As
an example, consider selection of Y - only

Change 9

3-46.9

TM 1-1520-237-10

mode. Press KYBD key two times. Observe that the GPS mode blinks. To enter
Y (for Y mode) press key LTR LEFT followed by key 9, or press key 9 only. A Y
will be displayed. Press ENT key. The entire display will blank out for less than one
second and the center display will now indicate: Y.
(5) Selection of DGNS mode of operation. As
an example, consider selection of GPS only mode of operation. Press KYBD key.
Observe that the DGNS mode blinks. To
enter G (for GPS mode) press key LTR
LEFT followed by key 3, or press key 3
only. A G will be displayed. Press ENT
key. The entire display will blank out for
less than one second and the DGNS mode
will now indicate: G (or * if GPS is not
available).
(6) Ground speed and ground track angle are
displayed on lines 3 and 4.
f. Procedure for entering/displaying present position or one of the 100 possible destinations in
MGRS. The DGNS has the capability to display 100 destinations (numbered 00-99).
100 destinations
-00 to 69

Standard waypoints.

-70 to 89

Data
load
only
waypoints,
observable but not changeable via
CDU keyboard. Used for national
airspace data such as VORs, NDBs,
and intersections.

-90 to 99

Target store waypoints (usable as
standard waypoints, but not as
route sequencing waypoints).

As an example, consider display of destination number
25.
(1) Enter datum as described in paragraph j.
below.
(2) Set MODE selector to MGRS.
(3) Set DISPLAY selector to WP/TGT.
(4) Notice the current destination number displayed. To display destination number 25
3-46.10

Change 9

press the INC or DEC key, or press key 2
then 5. This is a direct key entry action.
(5) Observe that the current destination
MGRS zone, area, and easting/northing
coordinates are now displayed. The destination number 25 and location name/ICAO
identifier also appears in the display.
(6) Entry for destination coordinates and location name/ICAO identifier: As an example,
consider entry of zone 18T, area WN, easting 5000, northing 6000, and ICAO identifier BANDO.
(7) To enter key board mode press the KYBD
key. Observe 9kybd9 displayed in the bottom right corner of the display. (Destination number blinks.) Press KYBD again.
(Zone field blinks.) To enter 18T press
keys 1, 8, LTR MID, 7.
(8) Press KYBD. (Area and northing/easting
blinks.) To enter WN5000 6000 press keys
LTR MID, 8,LTR MID, 5, KYBD, 5, 0,
0, 0, 6, 0, 0, 0.
(9) Press KYBD. (Location name/ICAO identifier blinks.) To enter BANDO press keys
LTR MID, 1, LTR LEFT, 1, LTR MID,
5, LTR LEFT, 2, LTR RIGHT, 5.
(10) To store the displayed information into the
selected destination display position press
the ENT key.
NOTE
To access P, press the LTR LEFT key followed by key 6. Another way to access P is
to display waypoint 99 then press the INC
key or display waypoint 00 then press the
DEC key.
Waypoints cannot be recalled by location
name/ICAO identifier.
g. Procedure for entering/displaying present position or one of the 100 possible destinations in
LAT/LONG. The DGNS set has the capability
to display 100 destinations (number 00-99).
100 destinations

TM 1-1520-237-10

-00 to 69

Standard waypoints

-70 to 89

Data
load
only
waypoints,
observable but not changeable via
CDU keyboard. Used for National
Airspace Data such as VORs,
NDBs, and intersections.

-90 to 99

Target store waypoints (Usable as
standard waypoints, but not as
route sequencing waypoints).

As an example, consider display of destination number
25.

(1) Enter the datum as described in paragraph
j. below.
(2) Set MODE selector to LAT/LONG.

(10) To store the displayed information into the
selected destination display position press
the ENT key. Display indicates:
N41° 10.13
E035° 50.27.
NOTE
To access P, press the LTR LEFT key followed by key 6. Another way to access P is
to display waypoint 99 then press the INC
key or display waypoint 00 then press the
DEC key.
Waypoints cannot be recalled by location
name/ICAO identifier.
h. Procedures for entering variation and landing
mode data.

(3) Set DISPLAY selector to WP/TGT.
(4) Notice that the current destination number
is displayed. To display destination number 25 press the INC or DEC key, or press
key 2 then 5. This is a direct key entry
action.
(5) Observe that the current latitude and longitude coordinates are now displayed. The
destination number 25 and location name/
ICAO identifier appears in the display.
(6) Entry of destination coordinates and location name/ICAO identifier: As an example,
consider entry of Latitude N41° 10.13 minutes and longitude E035° 50.27 minutes
and ICAO identifier BANDO.
(7) To enter keyboard mode press KYBD key.
Observe 9kybd9 display in the bottom right
corner of the display. (Destination number
blinks.) Press KYBD again. (Latitude field
blinks.) To enter N41° 10.13 press keys N,
4, 1, 1, 0, 1, 3.
(8) Press KYBD. (Longitude field blinks.) To
enter E035° 50.27 press keys E, 0, 3, 5, 5,
0, 2, 7.
(9) Press KYBD. (Location name/ICAO identifier blinks.) To enter BANDO press keys
LTR MID, 1, LTR LEFT, 1, LTR MID,
5, LTR LEFT, 2, LTR RIGHT, 5.

(1) Set MODE selector to MGRS positionaltitude entered/displayed in meters (LAT/
LONG may also be used-altitude entered/
displayed in feet).
(2) Set DISPLAY selector to WP/TGT position.
(3) Select the waypoint number desired by directly entering the two digit target number
or pressing the INC/DEC keys. Observe
the waypoint number entered and position
data.
(4) Press the ENT key and observe the waypoint number, variation and/or landing data
if entered.
(5) To enter a magnetic variation and/or landing mode data press the KYBD key to select the field for entry and enter the desired
data as shown in steps 6 through 10 below.
To end the entry operation press the ENT
key.
(6) Entry of variation: as an example, consider
entry of a variation of E001.2. Press keys
E, 0, 0, 1 and 2. The decimal point is inserted automatically. If no landing mode
data is to be entered, press ENT to complete the operation. Display indicates:
E001.2°.

Change 9

3-46.11

TM 1-1520-237-10

NOTE
An asterisk appearing in the variation field
indicates the variation is not entered. Variations may not be entered for waypoints containing target motion.
(7) The bottom two lines indicate the MSL altitude, desired glideslope, and the desired
inbound approach course (IAC) to the indicated destination. As an example, consider entry of a glideslope of 8° an IAC of
270°, and an altitude of 230 meters, for
destination number 25.
(8) Press the KYBD key to blink the glide
slope field. Enter glideslope. The maximum allowable glideslope is 9 degrees. In
the example enter 8 for an eight degree
glideslope.
(9) Press the KYBD key to blink the inbound
approach course field. Enter a three digit
inbound approach course angle. In the example enter 2, 7, 0 to enter a 270 degree
inbound approach course. Press the ENT
key to complete the operation.
(10) Press the KYBD key to blink the altitude
field. Press the INC/+ key to enter a positive altitude, press keys 2, 3, 0 (the leading
zeros may be omitted) for the altitude of
230 meters in the example.
i. Procedures for entering target motion and direction. In MGRS mode, target speed is entered in km/hr; in LAT/LONG mode, target
speed is entered in knots.
(1) Set the MODE selector to LAT/LONG
(MGRS may be used).
(2) Set the DISPLAY selector to WP/TGT
and select the target number desired (00-69
or 90-99) by directly entering the two digit
target number or INC/DEC keys. Observe
the waypoint number entered and position
data.

3-46.12

Change 10

(3) Press the ENT key and observe the waypoint number, variation and/or landing data
if entered.
(4) Press the ENT key and observe the target
speed and direction page.
(5) To select target speed press the KYBD key
twice and enter the target speed. The maximum target speed that may be entered is 50
knots.
(6) To select the target direction press the
KYBD key and enter the target direction.
(7) To end the entry operation press the ENT
key. In the example target 93 has a speed
of 18 knots and a bearing of 128 degrees.
At the time the ENT key is pressed and
released, the target position will begin to
be updated as a function of time based on
the speed and direction entered.
NOTE
To abort/cancel and entry of target motion,
enter a target speed of 000 using the above
procedure.
j. Procedure for entering/displaying datum (Table
3-1.1) or clearing all waypoints.
(1) Set the MODE selector to MGRS position
(LAT/LONG may also be used).
(2) Set the DISPLAY selector to DATUM/
ROUTE.
(3) To select the datum field press the KYBD
key.
(4) Entry of ellipsoid: as an example consider
entry of 47, the code of the WGS 84 datum. Press keys 4 and 7. Press the ENT
key, the display shall show DATUM: 47.

TM 1-1520-237-10

NOTE
Entering a new datum number to a particular
waypoint applies that datum to all waypoints
and converts their coordinates accordingly.
For example, assume that the datum for
waypoint 22 is datum 47 and the datum for
waypoint 23 is datum 25. The datum number
must be changed from 47 to 25 prior to entering data for waypoint 23. This will change
the displayed coordinates for waypoint 22
because they have been converted from datum 47 to datum 25. The actual ground position of waypoint 22 has not changed. Extreme care must be taken not to confuse
these newly converted coordinates with
those originally entered.

SEA CURRENT
SP:XXXKn
DIR:XXX°
(5) Entry of sea current speed and direction: as
an example, consider the entry of 4 knots
and 135 degrees. Press KYBD key. Observe that the speed field blinks.
(6) To enter speed, press keys 0, 0 and 4. The
speed indicates 004Kn. The maximum sea
current speed that may be entered is 50
knots.
(7) Press KYBD key. The direction display
blinks.

(5) To clear all waypoints, variations, landing
data and target motions, enter RDW for
the datum.

(8) To enter direction, press keys 1, 3, and 5.
Direction indicates 135°.

k. Procedure for entering sea current speed and
direction for water motion correction. (Required only if GPS is not available.)

(9) Press ENT key. The display momentarily
blinks and then reappears.

NOTE

Not required or necessary when in combined
or GPS mode. In MGRS mode, surface
wind speed is entered in km/hr; in LAT/
LONG mode, surface wind speed is entered
in knots. Leading zeros must be entered.
Wind direction is defined as the direction
from which the wind originates.

To abort entry of sea current, enter a sea
current speed of 000 using the above procedure.

(1) Set MODE selector to LAT/LONG
(MGRS may be used).

Table 3-1.1. Datums (AN/ASN-128B)

(2) Set DISPLAY selector to WIND-UTC/
DATA and observe the standard wind
speed and direction display.

NAME

ELLIPSOID
ID

NOTE

Only use map datums
WGS-84 and NAD-27. Other
map datums were not verified
using the Aviation Mission
Planning System (AMPS),
and should not be used.

(3) Press the ENT key twice to display the selection menu.
1>SEA CURRENT
2>SURFACE WIND

Adindan

3>GPS STATUS

ARC 1950

Australian Goodetic 1966

Bukit Rimpah

Camp Area Astro

4>DATA LOAD end
(4) Press the 1 key to select SEA CURRENT.
The display indicates:

Change 10

3-46.13

TM 1-1520-237-10

Table 3-1.1. Datums (AN/ASN-128B) (Cont)

Table 3-1.1. Datums (AN/ASN-128B) (Cont)
ID

NAME

ELLIPSOID
ID

NAME

Djakarta

(Corrego

European 1950

South American
Alegre)

Geodetic Datum 1949

South American (Campo Inchauspe)

Ghana

South American (Chua Astro)

Guam 1963

South American (Yacare)

G. Segara

Tananarive Observatory 1925

G. Serindung

Timbalai

Herat North

Tokyo

Hjorsey 1955

Voirol

Hu-tzu-shan

Indian

Special Datum, Indian Special

Ireland 1965

Special Datum, Luzon Special

Kertau (Malayan
Triangulation)

Special Datum, Tokyo Special

Special Datum, WGS 84 Special

Revised

ELLIPSOID
ID

Liberia 1964

USER ENTERED

Luzon

WGS72

Merchich

WGS84

Montjong Lowe

Nigeria (Minna)

North American 1927 (CONUS)

North American (Alaska and
Canada)

Old Hawiian, Maui

Old Hawiian, Oahu

Old Hawiian, Kauai

Ordance Survey of Great
Britain 1936

Qornoq

Sierra Leone 1960

South American (Provisional
1956)

3-46.14

Change 10

l. Procedure for entering surface wind speed and
direction for water motion correction.
(1) Set MODE selector to LAT/LONG
(MGRS may also be used).
(2) Set DISPLAY selector to WIND-UTC/
DATA and observe the wind speed/
direction display. Display will not appear
if TAS sensor is not installed.
(3) Press the ENT key twice to display the selection menu.
1>SEA CURRENT
2>SURFACE WIND
3>GPS STATUS

TM 1-1520-237-10

4>DATA LOAD end

time-to-go to the fly-to destination are
displayed.

(4) Press key 2 to select SURFACE WIND.
The display indicates:
SURFACE WIND
SP:XXXKn
DIR:XXX°
(5) Entry of wind speed and direction: as an
example, consider the entry of 20 knots and
150 degrees. Press KYBD key. Observe
that the wind speed field blinks.
(6) To enter the speed, press keys 0, 2, and 0.
The wind speed indicates 020. The maximum surface wind speed is 50 knots.
(7) Press KYBD key. The direction display
blinks.
(8) To enter direction, press keys 1, 5, and 0.
Wind direction indicates 150°.

(2) When the aircraft is over the destination,
press KYBD key. Observe that the display
freezes.
(3) Position update can be effected by pressing
the ENT key. The computer updated the
present position at the time the KYBD key
was pressed by using the stored destination
coordinates, and adding to them the distance traveled between the time the KYBD
key was pressed and the ENT key was
pressed. In addition, if an associated variation for the stored destination exists, the
present position variation is also updated.
(4) If a present position update is unnecessary
(as indicated by an appropriately small
value of DISTANCE to go on overflying
the destination), set the DISPLAY selector
to some other position - this action aborts
the update mode.
b. Updating present position from a landmark.

(9) Press ENT key. The display momentarily
blinks and then reappears.

Method 1 (Unexpected update)

NOTE

To abort entry of surface wind speed and
direction, enter a surface wind speed of 000
using the above procedure.

There are two methods for updating present
position from a landmark. Method 1 is particularly useful if the landmark comes up
unexpectedly and the operator needs time to
determine the coordinates. Method 2 is useful when a landmark update is anticipated.

3.17A.11 Flight Procedures.
NOTE

(1) Set DISPLAY selector to PP position.
This procedure is applicable to the Doppler
only mode. Present position is automatically
updated when DGNS is in combined mode.
a. Updating present position from a stored destination.
NOTE
The preface is: The aircraft is flying to a
destination. Destination is set to the number
of the desired destination.
(1) Set DISPLAY selector to DIST/BRG/
TIME position. Distance, bearing and

(2) Overfly landmark and press KYBD key.
The present position display shall freeze.
(3) Compare landmark coordinates with those
on display.
(4) If the difference warrants an update, enter
the landmark coordinates by pressing the
KYBD key to blink the field to be changed,
enter coordinates, then press the ENT key.
The computer updates the present position
(from the time the KYBD key was pressed)
to the landmark coordinates, and adds to
the updated present position the distance

Change 10

3-46.15

TM 1-1520-237-10

traveled between the time the KYBD key
was pressed and the ENT key was pressed.
(5) If an update is not desired, set the DISPLAY selector to some other position.
This action aborts the update mode.
Method 2 (Anticipated update)
(1) Set DISPLAY selector to WP/TGT position.
(2) Access P by pressing the LTR LEFT key
followed by key 6, entering destination 00
then pressing the DEC key, or entering
destination 99 then pressing the INC key.
(3) Press KYBD key. Observe that the display
freezes.
(4) Manually enter the landmark coordinates
by pressing the KYBD key to blink the
field to be changed and enter the coordinates.

(2) Set DISPLAY selector to XTK/TKE. Observe standard crosstrack (XTK) and track
angle error (TKE) display. (DIST/BRG/
TIME may also be used).
(3) To display Fly-To destination 43 press the
INC or DEC key, or press key 4 then 3.
This is a direct key entry action.
Left-Right Steering Signals
c. There are two methods the pilot may use to
fly-to destination, using left-right steering signals displayed on the computer-display unit. As
an aid to maintaining course, set DISPLAY selector to XTK/TKE position and steer vehicle
to keep track angle error (TKE) nominally zero.
Left-right steering signals may be used when
flying the shortest distance to destination from
present position (Method 1) or when flying a
ground track from start of leg to destination
(Method 2).
Method 1

(5) When overflying landmark, press ENT
key.
(6) If an update is not desired, set the DISPLAY selector to some other position.
This action aborts the update mode.
3.17A.12 Fly-To Destination Operation.
a. Initialization of Desired Course. When a fly-to
destination is selected such as at the start of a
leg, the present position at the time is stored in
the computer. A course is then computed between the selected point and the destination. If
the aircraft deviates from this desired course,
the lateral offset or crosstrack distance error is
computed. Distance and bearing to destination,
actual track angle, and track angle error correction are computed from resent position to destination. See Figure 3-18.2 for a graphic definition of these terms.
b. Procedure for selecting one of 100 possible
Fly-To destinations (Direct/Default Mode). The
Doppler/GPS navigation set has the capability
of selecting a fly-to destination from 100 destinations (number 00-99). As an example, consider selecting Fly-To destination number 43.
(1) Set MODE selector to MGRS (LAT/
LONG or GPS LDG may also be used).
3-46.16

Change 10

When flying shortest distance to destination from present
position, set DISPLAY selector to DIST/BRG/TIME position and steer vehicle to bearing displayed. If the display
indicates a L (left) TKE, the aircraft must be flown to the
left to zero the error and fly directly to the destination.
Method 2
When flying a ground track, set DISPLAY selector to
XTK/TKE position. Steer vehicle to obtain zero for
crosstrack error (XTK). If XTK is left (L), aircraft is to
right of the desired course and must be flown to the left to
regain the initial course. Select the course deviation bar by
pressing, then releasing the DPLR GPS lens on the HSI
MODE SEL panel.
d. Procedure to enter route-sequence to-to mode.
The Doppler/GPS navigation set has the capability to navigate a course set up between two
destinations. As an example, consider navigating onto a course starting from destination
number 62 and ending at destination number
45.
(1) Set MODE selector to MGRS (LAT/
LONG may also be used).
(2) Set DISPLAY selector to DATUM/
ROUTE.

TM 1-1520-237-10

MAGNETIC NORTH

TRUE NORTH
FLY−TO
DESTINATION

BEARING (B)
MAGNETIC NORTH

DISTANCE (D)

TKE
H

GROUND TRACK

T
XTK

AIRCRAFT AXIS

PRESENT POSITION
START
OF LEG

WIND SPEED
AND DIRECTION

LEGEND
XTK − CROSSTRACK DISTANCE ERROR
TKE − TRACK ANGLE ERROR
B − BEARING TO DESTINATION
T − ACTUAL TRACK ANGLE (TRUE TRACK)
D − DISTANCE FROM PRESENT POSITION
TO FLY−TO−DESTINATION
H − TRUE HEADING

AA9999
SA

Figure 3-18.2. Definition of Course Terms
(3) Press the ENT key. Observe that a menu
of special steering functions appears.
1>TO-TO
2>RANDOM
3>CONSEC
END

(4) To select the route-sequence to-to display
press key 1. Observe that TO-TO and selection mode appears in the display. The
display provides entry of starting and ending destination numbers.

(5) To enter keyboard mode press the KYBD
key. (START field blinks.) To enter starting destination 62 press keys 6, 2.
(6) Press KYBD key. (END field blinks.) To
enter ending destination 45 press keys 4, 5.
(7) Press KYBD key. (SELECT field blinks.)
Enter Y (yes) for mode selection. N (Default mode) may be entered to arm the system with the start and end destinations but
without entering the route-sequence to-to
mode, or to exit the Route-sequence to-to

Change 10

3-46.17

TM 1-1520-237-10

mode if the system is currently in that
mode. Then press the ENT key.
NOTE
There must be valid waypoint data to select
a waypoint as a starting or ending waypoint.
If not, upon pressing the ENT key, the invalid waypoint number will blink.
If an entry is changed after Y is entered for
selection, an N must be entered for the selection then it may be changed to Y. The
sequences must be flown from the beginning
waypoint. The route cannot be flown in reverse (R).
No target destination or destination with target motion may be included as to-to waypoints.
If the MODE switch is placed to the GPS
LDG position when TO-TO, RANDOM, or
RT SEQ CONSEC is selected, it will turn
off the route sequencing mode and change it
back to direct-to.
e. Procedure to enter route-sequence random
mode. The Doppler/GPS navigation set has the
capability to navigate through a sequence of
random number destinations. As an example,
consider navigating through destination numbers 32, 25, 74, 01, 48, 83, 35.
(1) Set MODE selector to MGRS (LAT/
LONG may also be used).

(5) Enter the sequence of destination numbers
by pressing the KYBD key to enter keyboard mode. (First destination field blinks.)
To enter destination 32 press keys 3, 2.
(6) Press KYBD key. (Next destination field
blinks.) Press keys 2, 5 to enter second
destination 25.
(7) Repeat step 6 until a maximum of ten destinations are entered or if less than ten need
to be entered, asterisks are left for remaining destinations.
(8) To complete the entry of the random sequence of waypoints press ENT key.
(9) To select the start field and enter the starting destination press KYBD key.
(10) To select the ending field and enter the
ending destination press KYBD key.
(11) Press KYBD key. (SELECT field blinks.)
Enter Y (yes) for mode selection. N (default) may be entered to arm the system but
without entering the route-sequence random mode, or to exit the Route-Sequence
Random mode if the system is currently in
that mode. An entry Y and R indicates a
choice of Y- flying in forward order, or Rflying in reverse order. To clear the random sequence, enter a C for selection.
Then press the ENT key.
NOTE

(2) Set DISPLAY selector to DATUM/
ROUTE.
(3) Press the ENT key. Observe that a menu
of special steering functions appears.
1>TO-TO

The sequence must be flown from the beginning waypoint.
No target destinations or destinations with
target motion may be included as route sequence random waypoints.

2>RANDOM
3>CONSEC
END
(4) To select the route-sequence random display press key 2. Observe that RT SEQ
RANDOM now appears in the display followed by the sequence of destination numbers and a continuation prompt.
3-46.18

Change 4

If the MODE switch is placed to the GPS
LDG position when TO-TO, RANDOM, or
RT SEQ CONSEC is selected, it will turn
off the route sequencing mode and change it
back to direct-to.
f. Procedure to enter route-sequence-consecutive
mode. The Doppler/GPS navigation set has the
capability to navigate through a sequence of

TM 1-1520-237-10

consecutively numbered destinations. As an example, consider navigating through destination
numbers 32 through 35.

RT SEQ CONSEC is selected, it will turn
off the route sequencing mode and change it
back to direct-to.

(1) Set MODE selector to MGRS (LAT/
LONG may also be used).

g. Procedure for displaying distance/bearing/time
information.

(2) Set DISPLAY selector to DATUM/
ROUTE.

(1) Set MODE selector to MGRS (LAT/
LONG or GPS LDG may also be used).

(3) Press the ENT key. Observe that a menu
of special steering functions appears.
1>TO-TO
2>RANDOM
3>CONSEC
END
(4) To select the route-sequence-consecutive
display press key 3. Observe that RT SEQ
CONSEC now appears in the display, followed by the starting and ending destination numbers, and mode selection.
(5) To enter keyboard mode, press the KYBD
key. (START field blinks.) To enter destination 32 press keys 3, 2.
(6) Press KYBD key. (END field blinks.)Press
keys 3, 5 to enter ending destination 35.
(7) Press KYBD key. (SELECT field blinks.)
Enter Y (yes) for mode selection. N (default mode) may be entered to arm the system but without entering the routesequence-consecutive mode, or to exit the
route-sequence-consecutive mode if the
system is currently in that mode. An entry
of Y and R indicates a choice of Y- flying
in the forward order, or R- flying in reverse order.
NOTE
The sequence must be flown from the beginning waypoint.
No target destinations or destinations with
target motion may be included as route sequence consecutive waypoints.
If the MODE switch is placed to the GPS
LDG position when TO-TO, RANDOM, or

(2) Set DISPLAY selector to DIST/BRG/
TIME.
(3) Observe that the distance-to-go in kilometers (to the fly-to destination), bearing, and
time-to-go appears on the bottom two lines
of the display. (Distance is in nautical
miles when MODE selector position is
LAT/LONG.) Bearing-to-destination is
displayed in degrees, and the time-to-go is
displayed in hours, minutes, and tenths of a
minute.
(4) The display of the second line depends on
the current steering mode as follows:
(a) Direct-To steering (default): Fly-to
destination number and ICAO identifier
are
displayed.
Example:
58:BANDO
(b) To-To Steering: TO-TO-:XX TO YY
where XX is the 9To-To9 start-of-leg
destination number, and YY is the
9To-To9 fly-to destination number.
(c) Route-sequence steering (both consecutive
and
random):
RTRANDOM:XX TO YY where XX is
the current route-sequence fly-to destination number, and YY is the next
destination number in the sequence.
Approximately 10 seconds before
overflying the fly-to destination, the
system automatically ’pickles’ to the
next destination, and the new fly-to
destination number blinks for 10 seconds then stops blinking.
h. Procedure for displaying present position and
GPS Altitude.
(1) Set the MODE selector to MGRS (LAT/
LONG or GPS LDG may also be used).
Set the DISPLAY selector to PP and observe present position display.

Change 2

3-46.19

TM 1-1520-237-10

(2) To display present position variation and
GPS altitude press the ENT key. Present
position variation may be entered by pressing the KYBD key to select the variation
field. A variation is entered and the ENT
key is pressed.
i. Target Store (TGT STR) Operation. Two
methods may be used for target store operation.
Method 1 is normally used when time is not
available to preplan a target store operation.
Method 2 is used when time is available and it
is desired to store a target in a specific location.
Method 1 (uses location 90-99)
(1) Press the TGT STR key while flying over
target.
(2) Present position and variation are automatically stored in the target destination location which was displayed in the target store
field immediately prior to pressing the
TGT STR key.
Method 2 (uses locations 00-69 and 90-99)
(1) Set MODE selector to MGRS or LAT/
LONG position, depending on coordinate
from desired.
(2) Set DISPLAY selector to WP/TGT position.
(3) To access P, press the LTR LEFT key
followed by key 6 . Another way to access
P, is to display waypoint 99 then press the
INC key or display waypoint 00 then press
the DEC key.
(4) Press KYBD key when overflying potential target. Observe that display freezes and
kybd is displayed in the bottom right corner of the display indicating keyboard
mode. The destination number is now under keyboard control indicated by a blinking field.
NOTE
Do not press ENT key while destination is
set to P.
(5) If it is desired to store the target, enter the
two digit destination number and press the
ENT key.
3-46.20

Change 4

(6) If it is not desired to store the target, set the
DISPLAY selector momentarily or permanently to another position.
j. Procedure for entering landing mode.
(1) Set the fly-to destination by setting the
DISPLAY selector to either XTK/TKE/
KEY or DIST/BRG/TIME. Directly enter
the two digit destination number or use the
INC or DEC keys.
(2) Set MODE selector to GPS LDG.
(3) The DISPLAY selector continues to function as before. To switch between metric
and English units, press the ENT key.
NOTE
In this mode, the DGNS provides real-time
landing guidance information to the HSI and
VSI indicators. To display course deviation
information on VSI and HSI, press then release the DPLR GPS button on the HSI/VSI
MODE SEL panel.
k. Procedure for transferring stored destination/
target data from one location to another. The
following procedure allows the operator to
transfer (copy) stored destination/target data
from one destination/target location to another
destination location. The transferred data consists of destination name/ICAO identifier, location, variation, and landing information. For illustrative purposes only, it is assumed that the
operator wants to put the coordinates of stored
target 97 into the location for destination 12.
(1) Set DISPLAY selector to WP/TGT position.
(2) Press key 9 then 7.
(3) Press KYBD key, press key 1 then 2.
NOTE
Location name/ICAO identifier, variation,
and landing data may be deleted by first displaying the waypoint, pressing the KYBD
key, then the ENT key.
(4) Press ENT key.

TM 1-1520-237-10

l. Operation during and after a power interruption. During a power interruption, the stored
destination and target data and present position
are retained by non-volatile RAM inside the
CDU. This makes it unnecessary to reenter any
navigation data when power returns. GPS satellite data are also retained by a battery inside
the SDC. This makes it unnecessary to reload
the crypto key or wait for the collection of any
almanac. Navigation will be interrupted during
the absence of power; however the present position will be updated when the GPS data becomes valid provided the DGNS mode has not
been selected as Doppler only. The pilot will
have to re-enter the GPS operating mode (M or
Y) using a single key (5 or 9). In the event the
CDU is initialized, the display will indicate
only EN when the CDU is operated. This is an
indication to the operator that previously stored
data has been lost and that spheroid/variation,
destinations, and calibration data must be entered. Present position needs to be entered only
if Doppler only mode has been selected. The
KYBD key must be pressed to clear the EN.
The pilot will have to re-enter the GPS operating mode (M only) using single key (5). The
computer initializes to the following: operating
mode to combined, present position variation to
E000.0, destinations and associated variations
to a nonentered state, wind speed (water motion) and sea current speed to 000, spheroid to
WGS 84 (WG-4), present position to N45°
00.00’E000°00.00’ (until updated by GPS), target store location to 90, along track calibration
correction to 00.0 percent, and magnetic compass deviation corrections to 000.0 degrees. The
following data must be entered:

(7) Enter each destination and its associated
variation.
m. Procedure for displaying aircraft heading, pitch,
and roll (Maintenance Function).
(1) Set the CDU mode switch to TEST and
observe the CDU test mode display.
(2) After the Doppler test is completed press
the ENT key.
(3) Observe the CDU display. The top three
display lines indicate, in degrees and tenths
of a degree, aircraft system heading, pitch,
and roll.
3.18 INTEGRATED INERTIAL NAVIGATION SYSTEM (IINS) AN/ASN-132(V). EH
a. The IINS is a self-contained integrated navigation
system capable of short and/or long-range missions which
can be updated whenever TACAN navigational facilities
exist or manually without TACAN data, and displays location of the helicopter on the control display unit (CDU).
The IINS consists of the following equipment:
TYPE
DESIGNATION

NAME

COMMON
NAME

C-11097/
ASN-132

Control
Indicator

Control
Unit

CV-3739/ASN132

Converter, Signal
Set

Signal Converter
Unit (SDC)

AN/UYK-64(V)2

Data Processing
Set

Navigation
Processor
Unit
(NPU)

RT-1159/A

ReceiverTransmitter,
Radio

TACAN RT

AN/ASN-141

Inertial
Navigation Set

Inertial
Navigation Unit
(INU)

MT-4915/A

Mounting Base,
Elect Equip

TACAN/SCU
Mount

(1) Press KYBD key.
(2) Set MODE selector to OFF momentarily,
to LAMP TEST for approximately one
second, and then to MGRS or LAT/
LONG.
(3) Select GPS M or Y mode.
(4) Select DGNS operating mode if other than
combined.

Display

(5) Enter datum.
(6) Enter present position if Doppler only has
been selected.

b. Auxiliary components of the IINS includes the SYSTEMS SELECT panel, INU blower assembly, INU battery assembly, and data bus couplers. The IINS provides

Change 8

3-46.21

TM 1-1520-237-10

accurate indications of the helicopter navigation parameters
including present position, velocity, altitude and heading
information. The system employs a serial data bus for data
interchange within the IINS and with external mission system computers. The IINS also interfaces with the helicopter
flight instruments and altimeter encoder. The multiplex data
bus system consisting of two buses (A and B), with only
one bus active at any given time. The other bus is in a
standby mode for redundancy purposes to provide a path
for data flow between the Standard Inertial Navigation System (STD INS), Signal Converter Unit, Navigation Proces-

3-46.22

Change 8

sor Unit, Control Display Unit, and the external mission
systems. Data to and from the TACAN receiver-transmitter
is first processed by the SCU before it is applied to the
multiplex data bus.
3.19 CONTROLS, DISPLAYS, AND FUNCTION.
The IINS controls and displays (Figure 3-19) are contained on the CDU. The function of each control is as follows:

TM 1-1520-237-10

KEY

CONTROL OR
INDICATOR

FUNCTION

Data display

Displays multiple flight parameters to the operator on seven
data lines and a scratch pad line on the face of the cathode
ray tube (CRT).

Line select keys

On both sides of data display lines 1, 3, 5, and 7 are pushbuttons (line select keys) which perform functions as defined by the legend adjacent to the key on the data display.
If a line select key is active on a particular page, an arrow
will appear in the character space closest to the key. Arrows will be oriented toward the legend, up, or toward the
key (away from the legend). These orientations (with examples) are defined as follows:
1. g legend s (TH 358.3 s). If the arrow points toward the
legend a numeric entry (entered into the scratch pad on line
8) is allowed by pressing the adjacent line select key.
2. s legend g (MAG g). If the arrow points away from the
legend, pressing the adjacent line select key will initiate the
function described by the legend. For example, pressing the
line select key adjacent to 9MAG g9 will change the display
to MAG Heading (MH) and MAG VAR (MV).
3. d legend d (d T/R). If the arrow points up, the legend
indicates current mode status and pressing the adjacent line
select key will change the mode. If no arrow appears next
to a legend then that line select key performs no operation.
4. Up or down pointing arrows on the sixth line of the data
display allows operator to slew display one page up or
down by pressing page slew toggle switch up or down.

Alphanumeric keys

Alphanumeric entries are made, by pressing one of the ten
character keys on the keyboard and will appear first in the
scratch pad (line 8). Each actuation of a key will cause a
character to be displayed from left to right in the scratch
pad. When using multiple letter keys (e.g., KLM/5), letters
K, L, or M can be entered into the scratch pad by successive actuations of the KLM/5 key. The 0-9 and. keys shall
enter the respective number of decimal point unless the
keyboard is in the letter mode. When the LTR/USE key is
pressed, LTR is annunciated to the right of the scratch pad,
and the next keystroke will enter an alpha character. When
the desired data appears in the scratch pad, it will be entered by pressing the line select key adjacent to the data
being updated. When the line select key is pressed, the
scratch pad contents will be checked for proper range and
format. If the entry is valid, it will be transferred to the
IINS, read back, and displayed adjacent to the line select
key. Completion of this cycle will clear the scratch pad.

CLR key

Used for erasing scratch pad parameters before entry. First
actuation clears the last number or letter entered, second
actuation clears the entire entry.

3-47

TM 1-1520-237-10

KEY

CONTROL OR
INDICATOR

FUNCTION

BRT control

Controls brightness of the data display from full on to full
off.

0 key

Used to enter number 0 into the scratch pad.

-/v key

Used to enter a minus symbol or decimal into the scratch
pad. When pressed, v will be entered into the scratch pad.
When LTR/USE key is pressed, then -/v key is pressed, will be entered into the scratch pad. To use the - in the
scratch pad, the LTR/USE key must be pressed again.

LTR/USE key

When pressed, allows letters to be entered into the scratch
pad. When pressed a second time, signals the CDU to use
the character that was just entered, and deletes LTR entry
mode.

FACK key

When pressed, signals the system that an annunciated failure has been recognized by the operator, and causes the
flashing annunciation to go to a steady annunciation.

Page select switch

Selects the type of information to be displayed. The following five categories of display pages can be selected:
NOTE

All CDU distance and speed displays in L/L mode are in
nautical miles (NM). All distance and speed displays in
UTM mode are in kilometers.
1. POS. Provides present position; universal transverse
mercator (UTM) or latitude/longitude (L/L) selection; magnetic heading selection; magnetic variation; true or magnetic heading; ground track; and ground speed.
2. INS. Provides inertial alignment status; barometric pressure; altitude; data zeroize; and access to system data and
unit tests.
3. DEST. Provides selected course entry; destination coordinates; UTM or L/L selection; range, bearing, and time to
destination; cardinal heading/distance.
4. STR. Provides selected course; range, bearing and time
to steerpoint; present position; UTM or L/L selection; cardinal heading/distance.
5. TCN. Provides both TACAN control and station data.
The TACAN control page provides power control; mode
selection; slant range and bearing to station; update enable;
and access to station pages. The TACAN station pages provide station magnetic variation; coordinates, channel; slant
range/bearing; and elevation.
11

3-48

Mode select switch

Selects eight different modes of operation for the IINS. The
mode select switch selects the following IINS modes of
operation:

TM 1-1520-237-10

KEY

CONTROL OR
INDICATOR
Mode select switch

FUNCTION
1. OFF. Turns off the IINS (removes power from TACAN
RT, STD INS, and CDU).
2. FAST. In this position the STD INS either performs a
stored heading alignment or best available true heading
(BATH) alignment. If a BATH alignment is performed,
true or magnetic heading information must be entered not
later than 1 minute after selecting the FAST mode. If magnetic heading information is not entered, system will assume a stored heading. After heading information is entered, present position may be entered if desired. FAST
alignment is a degraded mode of operation and should not
be used under normal conditions.
3. NORM. In this position the STD INS performs a gyrocompass alignment. Present coordinates must be entered
not later than 2 minutes after selecting the NORM mode.
4. NAV. This is the STD INS primary flight mode of operation. NAV is entered after satisfactory alignment conditions have been met.
5. UPDT. In this position the NPU freezes present position
data for a later manual position update by overflying a
known position designated by a mark.
6. ATTD. In this position the STD INS initiates an attitude
reference mode of operation. Although navigation processing is discontinued, the STD INS continues to provide a
stable reference frame for generation of roll, pitch, and inertial heading angles.
7. CAL. In this position the STD INS performs an automatic calibration of the gyro biases.
8. TEST. In this position the STD INS performs functional
performance tests, fault detection, and fault localization
checks.

DEST switch

Three position toggle switch used to increment/decrement
selected destination. The number of the selected destination
appears on line 1 of the data display. Up increases and
down decreases the selected destination.

BIT indicator

Used to indicate the results of all internal CDU tests. White
indicates a failure and black indicates test passed.

STR switch

Three position toggle switch used to increment/decrement
selected steer point. The number of the selected steerpoint
appears on line 1 of the data display. Up increases and
down decreases the steerpoint number.

MRK key

Used to signal the STD INS to note the current position and
use it for one of two of the following purposes:

3-49

TM 1-1520-237-10

KEY

CONTROL OR
INDICATOR

FUNCTION

MRK key

1. Store as a markpoint (destination A thru F) when the
mode select switch is in the NAV position.
2. Store present position relative to selected destination for
possible updating when mode select switch is in the UPDT
position.

Page slew switch

Three position toggle switch used to slew data display one
page up or down by pressing page slew switch up or down.
3.22 SYSTEMS SELECT PANEL.

3.20 SIGNAL CONVERTER UNIT,
CV-3739/ASN-132. EH
The SCU performs data processing to convert the
TACAN RT Aeronautical Radio Incorporated (ARINC) inputs and outputs to corresponding serial data formats for
transmission over the multiplex data buses to the NPU and
CDU. The SCU can communicate via one of the two multiplex data buses. Although the SCU communicates over
only one multiplex data bus at a time, it can monitor both
buses continuously to determine over which bus valid data
communications are taking place. Redundant portions of
the SCU circuitry are isolated to ensure that a failure of one
bus does not degrade performance of the remaining bus.
3.21 TACAN NAVIGATIONAL
TRANSMITTER, RT-1159/A. EH

SET

3-50

The SYSTEMS SELECT panel (Figure 3-20) consists
of two switch light indicators, located on the center lower
edge of the instrument panel. It provides a switching capability for utilization of IINS through a relay assembly. The
SYSTEMS SELECT panel operates as follows:

HDG
DG:

RECEIVER-

The position error of an inertial navigation system increases with time, therefore, a position reference sensor is
used to update the inertial data, and thereby bound the
time-growing position error. The IINS derives position updates from the TACAN RT range and bearing measurements. The TACAN RT determines the relative bearing and
range of the helicopter from a selected TACAN ground
station. The TACAN RT operates within 390 nautical miles
of a TACAN ground station. Since the TACAN system
operating limit is line of sight, the actual operating range is
dependent on helicopter altitude. The TACAN system operates on a selected channel from 252 available channels.
The 252 channels are equally divided into 126 x-channels
and 126 y-channels with both x- and y-channels spaced at 1
MHz intervals. Upon being interrogated by the TACAN
RT, the ground station beacon transmits a signal. From the
return signal, the TACAN RT computes bearing and distance values for updating the inertial system information.
The TACAN RT outputs are processed by the SCU for
compatibility with the multiplex data buses. The TACAN
RT also produces and transmits distance information when
interrogated in the air-to-air operation another TACAN
equipped aircraft, however, this air-to-air mode precludes
using the TACAN information to update the IINS.

IINS:

ASN-43 directional gyro output is
displayed on the HSI’s. ASN-43
interface with the VSI/HSI Mode
Select System, the SAS/FPS flight
computer, the civil navigation
system,
and
the
Command
Instrument System (CIS).

IINS heading output is displayed
on the HSI’s. IINS interface with
the above system, replacing the
ASN-43.

ATT
VG:

CN-1314 Pilot or copilot vertical
displacement gyro output is
displayed on respective VSI’s and
used by the SAS/FPS computer as
determined by the VSI/HSI MODE
SEL VERT GYRO setting.

IINS:

INU output is displayed on the
VSI’s and is used by the SAS/FPS
computer depending on the VSI/
HSI MODE SEL VERT GYRO
setting.

TM 1-1520-237-10

ABC
MRK

STR

I
I
N
S

INS

DEST

STR
TCN

POS

BIT

NORM
DEST

NAV

UPDT

FAST

DEF
3

GHJ

KLM

NPQ

RST

UVW

XYZ

CLR

LTR
USE

OFF

BRT

TEST

FACK

CAL

ATTD

AA0391
SA

Figure 3-19. CDU Controls and Indicators

NOTE

If the IINS is to be turned OFF
during flight, the IINS should be
deselected on the SYSTEMS
SELECT panel prior to IINS turn
OFF.

3.23 PILOT AND COPILOT VSI/HSI MODE SEL
PANEL. EH
The VSI/HSI MODE SEL Panel (Figure 3-21) modified
for IINS, operates the same as the UH-60A. The IINS
switch operation is as follows:
IINS

3-51

TM 1-1520-237-10

MODE SEL

SYSTEMS SELECT
VOR
ILS

BACK
CRS

FM
HOME

IINS

VOR
ILS

BACK
CRS

FM
HOME

NORM
ALTR

PLT
CPLT

NORM
ALTR

ADF
VOR

TURN
RATE

CRS
HDG

VERT
GYRO

BRG
2

IINS
DG

IINS

HDG

ATT

AA0392
SA

Figure 3-20. SYSTEMS SELECT Panel

Selection of IINS will display IINS calculated
range, bearing, and course deviation to the
steerpoint on the associated HSI. Range is displayed as distance (KM), bearing by the #1
pointer, and deviation by the course deviation
bar.
Selection of IINS disconnects the VOR (ARN123) TO/FROM output to the HSI’s and connects the SCU TO/FROM output to the HSI’s.
To select IINS on the MODE SEL panels,
IINS must be selected on the SYSTEMS SELECT panel. Also the CDU must be on and in
the NAV mode.
3.23.1 Valid Entry Procedures. The following paragraphs describe valid entry formats for data which may be
entered on each of five main pages. An inward pointing
arrow indicates that data can be entered on that line by
pressing the adjacent line select key.
a. POS (Position) Page. The position page provides for
entry of mag/true heading (BATH alignment), present position (FAST/NORM alignment) and magnetic variation.
(1) Magnetic True Heading Entry. Magnetic heading
and magnetic variation or true heading may be entered during the first 60 seconds of a FAST alignment. Scratch pad
entries may be up to four numeric digits including an optional decimal point. If no decimal point is entered, whole
degrees are assumed. Leading zeros are optional.
(2) Latitude Entry. Key in N or S and then the numeric
digits. The first two digits are degrees, third and fourth are

3-52

AA0393
SA

Figure 3-21. HSI/VSI MODE SEL Panel

minutes and fifth and sixth are seconds. A leading zero
must be entered for any latitude less than 10 degrees. Entry
examples:
PREVIOUS
VALUE

SCRATCH
PAD
CONTENTS

ENTERED
VALUE

S 6° 15 34

N263415

N 26° 34 15

N 33° 25 15

2634

N 26° 34 00

S 46° 13 00

S26

S 26° 00 00

S 46° 13 00

S 26° 00 00

(3) Longitude Entry. Key in E or W and then the numeric digits. A leading zero must be entered for a longitude
less than 100 degrees and two leading zeros for a longitude
less than 10 degrees.
PREVIOUS
VALUE

SCRATCH
PAD
CONTENTS

ENTERED
VALUE

E 176° 16 00

W1263415

W 126° 34 15

TM 1-1520-237-10

Table 3-2. Spheroid Data Codes

(Cont)

PREVIOUS
VALUE

SCRATCH
PAD
CONTENTS

ENTERED
VALUE

CODE

MODEL

ABBR

E 176° 16 00

12634

E 126° 34 00

Airy

ARY

W 135° 42 32

E126

E 126° 00 00

Hough

HGH

E 120° 16 24

126

E 126° 00 00

South American

SAM

Modified Everest

MEV

WGS.72

WGS

(4) Spheroid or Grid Zone Entry. Either spheroid or
grid zone may be entered. Spheroid entries consist of numbers 0 through 10 and are an alpha display as listed in Table
3-2. Grid zone entries consist of two numbers and alpha
character. Entry examples.

PREVIOUS
VALUE

SCRATCH
PAD
CONTENTS

ENTERED
VALUE

16T INT

16T CL6

16T INT

18T

18T INT

16T INT

18T1

18T CL6

(5) Area/Eastings/Northings Entry. Scratch pad entries
may be made for area, eastings and northings, just area, or
just eastings and northings. Entries for area must consist of
two alpha characters. Entries for eastings/northings must be
2, 4, 6, 8, or 10 digits. Digits will be evenly split between
eastings and northings with trailing zeros inserted. Although entries may be made and sent to the INU to a resolution of 1 meter, the display will round to the nearest 10
meters. The following illustrates several examples:
Table 3-2. Spheroid Data Codes

NOTE
When mission equipment operator selects
WGS 1984, CDU is code 10.

PREVIOUS
VALUE

SCRATCH
PAD
CONTENTS

ENTERED
VALUE

AU 1234 5678

UV23456789

UV 2345 6789

AU 1234 5678

23456789

AU 2345 6789

AU 1234 5678

2345678

AU 2340 6780

AU 1234 5678

2367

AU 2300 6700

AU 1234 5678

AU 2000 6000

AU 1234 5678

UV 1234 5678

(6) Magnetic Variation Entry. Scratch pad entries consist of an E/W and up to four numeric digits including
decimal point. If no decimal point is entered, while degrees
are assumed. The range of entries is 0.0° to E/W 180.0°.
For entries greater than or equal to 100°, only whole degrees are displayed. The following gives some entry examples.

CODE

MODEL

ABBR

SCRATCH PAD
CONTENTS

ENTERED VALUE

International

INT

E20

Clark 1866

CL6

W10.9

Clark 1880

CL0

E.7

E0.7

Everest

EVR

Bessel

BSL

Australian National

AUS

b. INS (Inertial) Page. The INS page provides miscellaneous control/display functions such as entry of altitude
and barometric pressure and provides access to INU and
NPU memory.
Change 5

3-53

TM 1-1520-237-10

IN0

CL6
IN0

BE0
CL6
EV0

CL6

IN0

CL0
BE0
AU0
IN0

IN0

AA8669A
SA

Figure 3-22. Doppler World UTM Spheroids (AN/ASN-128)
(1) Manually Entered Altitude (MALT). Field altitude
must be entered to the nearest 100 ft. MSL during alignment; however, manually entered altitude may be entered
any time during the mission to override barometric altimeter. The range of valid entries is from -1000 to +65,520
feet in increments of 100 feet. Entries shall delete MALT
by causing an output of -65,520 feet over the barometric
pressure.

SCRATCH PAD
CONTENTS

ENTERED VALUE

2.0

3-54

Change 5

SCRATCH PAD
CONTENTS

ENTERED VALUE

10.0

10.5

(2) Barometric Pressure (BARO). Barometric pressure
must be entered (0.01 in Hg) during alignment. The information is used by the NPU to initialize the scale factor of
encoding altimeter data during alignment.
(3) DATA Page.

TM 1-1520-237-10

(a) Press the line select key adjacent to line 5 right
(DATA).
(b) Line 7 of the DATA page provides the capability
to enter and read the contents of various INU registers.

Although the CDU will accept entered memory addresses
with a range of 0 to 65,535 (decimal), the INU will not
accept all of these as valid. If an illegal address is entered,
the illegal address and the message 9ENTRY REJECTED9
will alternately appear in the scratch pad. Pressing the CLR

Change 5

3-54.1/(3-54.2 Blank)

TM 1-1520-237-10

key will clear the scratch pad. Register contents that are
entered may be any six alphanumeric characters plus sign.
Many of the INU registers are 9read only9. That is, their
contents can be read but not altered. If an attempt is made
to change the contents of one of these registers, 9ENTRY
REJECTED9 will appear as described above.
c. DEST (Destination) Page. Two types of data may be
entered on the destination page, destination coordinates,
and course to destination.
(1) Destination Coordinates Entry. Destination coordinates may be entered during any phase of the mission. Either LAT/LONG or MGRS (UTM) coordinates may be entered. Coordinate selection is provided on line 7 (display
right).
(a) Latitude entry described in paragraph 3.23.1.a.2..
(b) Longitude
3.23.1.a.3..

entry

described

paragraph

P O S I T I O N

T H

3 5 8 . 3

M V = E 1 0

M H

3 4 8 . 0

G T K

. 3

3 5 9 . 6

1 3 T

I N T

W D
U V

G S

EXAM

1 2 3 4

1 8 1 . 2

P L EG R

3 4 5

1 2 3 4

0 2 5

U T M

[

]

NOTE
TO SELECT THIS PAGE, SET CDU PAGE
SELECT SWITCH TO POS AND PRESS
LINE 5 AND / OR LINE 7 RIGHT LINE
SELECT KEYS AS REQUIRED.

(c) Spheroid and zone entry described in paragraph
3.23.1.a.4..
(d) Area/Eastings/Northings entries described in
paragraph 3.23.1.a.5..

I D

AA0396
SA

Figure 3-23. Position Page

(2) Course to Destination Entry. The desired true
course to destination may be entered for each destination
during any phase of the mission.

(2) TACAN Station Page Entries. Parameter entries
are station location magnetic variation channel and elevation.

(a) Enter the true course in the scratch pad. Entries
may be up to four numeric digits with an optional decimal
point. If no decimal point is entered, whole degrees are
assumed. Leading zeros are optional.

(a) Latitude entry described in paragraph 3.23.1.a.2..

(b) Press line 1 right line select key. The true course
to destination will be displayed on line 1 right.
(c) System will utilize any previous course data. If
no data has been entered, the system will assume a true
course of 000.0 degrees.
d. STR (Steer) Page. This page contains no enterable
parameter.
e. TCN (TACAN) Pages. The TACAN pages consist of
the TACAN control page and TACAN station pages.
(1) TACAN Control Page Data Entry. The only data
entry on this page is channel number. Paragraph 3.23.3.f.
describes the channel number entry.

(b) Longitude
3.23.1.a.3..

entry

decribed

paragraph

(c) Spheroid and zone entry described in paragraph
3.23.1.a.4..
(d) Area/Eastings/Northings entries described in
paragraph 3.23.1.a.5..
(e) Station Magnetics Variation entry described in
paragraph 3.23.1.a.6..
(3) Station Channel Entry. The TACAN Station Pages
provide the capability to enter station channel for each of
the 16 stations. A total of 252 channel are possible (126 9X9
channels and 126 9Y9 channels). Unless 9Y9 is entered, an
9X9 channel is assumed.

3-55

TM 1-1520-237-10

3.23.2 Starting
MENT).

Procedure

(NORMAL

ALIGNCAUTION

1. Ensure that the following circuit breakers are
in:
a. NO. 1 AC PRI BUS circuit breaker panel
(1) INS
(2) XFMR PWR
(3) INU BATT PWR

Wait two minutes before returning mode
select switch to NORM. Failure to do so
will damage the INU.
b. If mode select switch was turned off, rotate
mode select switch to NORM. If an annunciation is still flashing, make an entry
on DA Form 2408-13-1. Refer to paragraph 3.23.5 for an explanation of annunciations.

(4) 26 VAC EQUIP PWR
3. Set page selector switch to POS (Figure 3-23).
b. NO. 1 DC PRI BUS circuit breaker panel
(1) CPLT ALTM
(2) IINS
c. NO. 2 AC PRI BUS circuit breaker panel
(1) TACAN
d. NO. 2 DC PRI BUS circuit breaker panel
(1) TACAN

NOTE
If UTM coordinates are selected, the COMPLETE UTM coordinates must be entered
for present position: GRID ZONE, SPHEROID, AREA, EASTINGS, and NORTHINGS.
4. Verify line 7 on right side display indicates desired COORDINATE SYSTEM (UTM or L/L).
If not press the line select key once to switch to
the desired coordinate system.
a. Enter GRID ZONE/SPHEROID or LATITUDE in scratch pad.

NOTE
b. Press line select key 5 left.
Present position must be entered during the
first two minutes of NORM alignment. If
present position is displayed, it must be reentered. A steady NAVRDY indicates INU
attitude data and degraded NAV performance are available. After turn-on, flashing
NAVRDY will be displayed on line 6 indicating full alignment.
2. To turn system on, set mode select switch to
NORM. CDU display will remain blank for 30
seconds after turn-on. If the CDU does not light
after 30 seconds, rotate the brightness control
on the CDU clockwise to provide a comfortable
intensity level.
a. Check for annunciations on line 2 of the
display. If any annunciation is flashing, return mode select switch to OFF.
3-56

Change 10

c. Enter AREA, EASTINGS, NORTHINGS,
or LONGITUDE in scratch pad.
d. Press line select key 7 left.
NOTE
If INU computed MV is changed, updated
MV will have to be manually entered as MV
lines are crossed. If INU computed MV is
utilized, automatic MV updating will be performed by the INU.
5. Verify line 3 on left side display indicates correct magnetic variation, MV.
a. If incorrect, enter MV in scratch pad.
b. Verify scratch pad entry is correct.

TM 1-1520-237-10

c. Press line select key 3 left.
d. Verify line 3 left displays: - > MV = XNN.
N. (The 9=9 sign indicates that a manual
MV was entered and automatic MV updating will not occur.)

I N S

D 2

I A L T

S 3

2 7 . 5

EXA

6. Rotate page select switch to INS (Figure 3-24).

T E S T S

MPL

2 9 . 0 1

a. Enter barometric pressure of present position in scratch pad.

B A R O

b. Press line select key 5 left.

1 9 . 9 M I N

c. Enter altitude of present position in scratch
pad (e.g., 156 ft is entered as 0.156 and
displayed as 0.2).

Z E R O I Z E

L A S T

D A T A
M R K

S T A T = A + H

[

NOTE

d. Press line select key 3 left.

TO SELECT THIS PAGE, SET CDU PAGE
SELECT SWITCH TO INS.

7. Rotate page select switch to DEST.
a. Press DEST toggle switch to select DEST
desired page.
b. Press line select key 7 right to display desired coordinate system (UTM or L/L).
c. Enter grid zone and spheroid or latitude in
scratch pad.

AA0394
SA

Figure 3-24. INS Page

ately press line select key 3 left until the
display shows REC.

d. Press line select key 5 left.

a. Turn ON TACAN by pressing line select
key 1 left.

e. Enter Area/Eastings/Northings or longitude
in scratch pad.

b. Press line 3 left until REC is displayed on
the CRT.

f. Press line select key 7 left.

c. Press page slew toggle switch to display
TACAN station zero page.

g. Press DEST toggle switch (Figure 3-19) to
increment to the next page.
8. Rotate page select switch to TCN.

(1) Enter magnetic variation in scratch
pad.
(2) Press line select key 3 left.
(3) Enter latitude in scratch pad.

WARNING
(4) Press line select key 5 left.
Potential radiation hazard exists at the
TACAN antenna when the TACAN is
turned on. Make sure that no person is
within 3 feet of antenna. When TACAN is
first turned on and line 3 left of CDU displays anything other than REC, immedi-

(5) Enter longitude in scratch pad.
(6) Press line select key 7 left.
(7) Press line select key 1 right to display
ACT.

3-57

TM 1-1520-237-10

(8) Enter channel number in scratch pad.
(9) Press line select key 3 right.
(10) Enter elevation of TACAN station on
scratch pad.
(11) Press line select key 5 right.
(12) Press page slew toggle switch to display next TACAN page.

b. Press ATT switch, INNS illuminates and
inertial derived pitch and roll is displayed
on the VSI.
3.23.3 Starting Procedure (FAST Alignment).
Switching to FAST mode commands the INU to perform
either a stored heading alignment or best available time
heading (BATH) alignment.
a. Stored Alignment.
NOTE

(13) Enter data as described in steps (a)
through (m).
NOTE
In order for HSI steering command to be
correct, a valid destination and steer point
must be entered prior to switching the
MODE selector to NAV.
Example: If DX is homebase (alignment
point), the SX STEER point is invalid as an
initial Destination/Steer Point.
9. Select an appropriate destination number and
toggle the STR toggle switch to indicate the
number, i.e., S1. (It is not necessary that the
DX and SX numbers agree, only that SX is the
desired destination.)
10. Set mode select switch to NAV. (Pull switch
up; then rotate.)
11. Set page select switch to TCN.
a. Press line select key 3 left to display T/R.
b. Press line select key 5 left to display
UPDT.
12. On HSI/VSI MODE SEL panel (Figure 3-21),
press IINS switch. Note that bearing to destination (No. 1 needle), range to destination,
course deviation and TO/FROM flag are displayed on the HSI.

3-58

CDU display will remain blank for 30 seconds after turn on. Barometric pressure must
be entered during alignment. Alignment will
be complete when data display line 6
NAVRDY indicator begins to flash if a normal alignment was performed and the mode
select switch was not set to NAV. Alignment will be complete when data display
line 6 NAVRDY indicator lights if a normal
alignment was performed and the mode select switch was set to NAV.
(1) Ensure system preoperational checks have
been performed and that aircraft power is
on.
(2) Set mode select switch (Figure 3-19) to
FAST.
(3) Set page select switch to POS.
(4) Observe that data display line 7 right indicates desired coordinate system (UTM or
L/L). If it does not, press line select key 7
right until the desired coordinate system is
deployed.
(5) If data display line 8 right indicates LTR,
press LTR/USE key to place the keyboard
in the numeric mode.

13. On SYSTEMS SELECT panel (Figure 3-20),
set switches and observe indications as follows:

(6) Observe that data display line 5 left and
line 7 left indicate present position latitude
and longitude or grid zone, spheroid area,
eastings and northings, respectively. If not,
normal or BATH alignment must be performed:

a. Press HDG switch, INS illuminates and inertial derived heading is displayed on the
HSI.

(7) Set page select switch to INS. Observe that
data display line 3 left indicates present
position altitude. If not, a change must be

TM 1-1520-237-10

made within the first 60 seconds of this
alignment.

(1) Ensure system preoperational checks have
been performed and that aircraft power is
on.

NOTE
(2) Set mode select switch to FAST.
The following steps are an example of entering barometric pressure. Substitute your own
barometric pressure when performing these
steps. Enter local barometric pressure to the
nearest 0.01 inches Hg.
When making keyboard entries, if an incorrect key is pressed, press CLR key as required and begin again.
(8) Enter local barometric pressure on data
display line 8 by pressing in sequence
ABC/N2, XYZ/ 9, -/v, 0, and 1 keys. Observe that data display line 8 indicates
29.01.
(9) Press data display line 5 left line select key.
Observe that data display line 5 left indicates - > BARO 29.01.
(10) Observe that data display line 7 indicates
alignment and status.
NOTE
Data display line 6 left indicates a flashing
NAVRDY if a normal alignment was performed and the mode select switch was not
set to NAV.
(11) When data display line 6 left NAVRDY
indicator lights, set mode select switch to
NAV.
b. BATH Alignment.
NOTE
CDU display will remain blank for 30 seconds after turn on. True or magnetic heading
must be entered during the first 60 seconds
of turn on. Present position must be entered
within 2 minutes of turn on. Barometric
pressure and altitude must be entered during
alignment.
Alignment will be complete when data display line 6 NAVRDY indicator lights.

(3) Set page select switch to POS.
(4) Observe that data display line 7 right indicates desired coordinate system (UTM or
L/L). If it does not, press line select key 7
right until desired coordinate system is displayed.
(5) If data display line 8 right indicates LTR,
press LTR/USE key to place the keyboard
in the numeric mode.
NOTE
The following steps are examples of entering present position data. Substitute your
own present position and heading when performing these steps. Either true heading or
magnetic heading can be entered. Magnetic
heading is entered by pressing line select
key. The following example uses true heading. When making keyboard entries, if an
incorrect key is pressed, press CLR key as
required and begin again.
(6) Enter true heading on data display line 8
by pressing in sequence DEF/3, KLM/5,
UVW/S8, -/ and DEF/3 keys. Observe that
data display line 8 indicates 358.3.
(7) Press data display line 1 right line select
key. Observe that data display line 1 right
indicates TH 358.3 ,-.
(8) If required, enter present position latitude
(or UTM, GRID ZONE and SPHEROID)
on data display line 8 by pressing in sequence LTR/USE, ABC/ N2, LTR/USE,
DEF/3, GHJ/W4, 1, 0, DEF/3, and 0 keys.
Observe that data display line 8 indicates
N341030.
(9) Press data display 5 left line select key.
Observe that data display line 5 left indicates - > N34° 10 30.
(10) If required, enter present position longitude
(or UTM area, EASTING and NORTHING) on data display line 8 by pressing in

3-59

TM 1-1520-237-10

T A C A N

T /

O N

S R N G

S T A

LE
AMP C

4 2 . 5

U P D T

O F

(14) Press data display line 3 left line select key.
Observe that data display line 3 left indicates -> AALT 8.6.

NOTE
H

B R G

1 2 5 X
1 2 2

Enter local barometric pressure to the nearest 0.01 inches Hg.

(15) Enter local barometric pressure on data
display line 8 by pressing in sequence
ABC/N2, XYZ/ 9, -/v, 0, and 1 keys. Observe that data display line 8 indicates
29.01.

1 0

S T A
3 S I G =

[

]

(16) Press data display line 5 left line select key.
Observe that data display line 5 left indicates - > BARO 29.01.
NOTE

(17) Observe that data display line 7 indicates
alignment time and status.

TO SELECT THIS PAGE, SET CDU PAGE
SELECT SWITCH TO TCN.

AA0395
SA

Figure 3-25. TACAN Control Page

sequence LTR/USE, GHJ/W4, LTR/USE,
1, 1, UVW/S8, DEF/3, and 0 keys. Observe that data display line 8 indicates
W1183530.
(11) Press data display line 7 left line select key.
Observe that data display 7 left indicates > W118° 35 30.
(12) Set page select switch to INS.
NOTE
When entering present position altitude, altitude must be entered to the nearest 100 feet
(mean sea level). The range of valid entries
from -1000 to 65,520 feet in increments of
100 feet. Entries are made in thousands of
feet.
(13) If required, enter present position altitude
on data display line 8 by pressing in sequence UVW/S8, -/v, and NPQ/E6 keys.
Observe that data display indicates 8.6.
This represents an altitude of 8,600 feet.
3-60

(18) When data display line 6 left NAVRDY
indicator lights, set mode select switch to
NAV.
3.23.4 In-flight Procedures.
a. MARK Operation. Current aircraft position may be
stored in one of the markpoint locations (destinations A-F)
by pressing the MRK key when in NAV mode. The location where present position was stored is displayed in the
CDU scratch pad regardless of currently selected page. Figure 3-24 illustrates 9MARK C9 in the scratch pad with the
STR page selected.
(1) The MARK locations are used in sequence (-A,
B, C, D, E, F, A, B,..). The MARK display will
remain in the scratch pad unless it is cleared
with the CLR key or the scratch pad is used to
enter some other data.
(2) Pressing the MRK pushbutton will freeze, for
30 seconds, the display of present position on
the Steering and Position pages; and the display
of cardinal headings/distance on both the destination and Steering Page. After 30 seconds or
after the CLR key is pressed, the current position will return.
b. Manual Updating (Overfly Position Updating). An
overfly update represents a manual position update technique in which the pilot overflies his selected destination
and signals the INU by pressing the MRK key. To initiate
a manual update, proceed as follows.

TM 1-1520-237-10

NOTE

(a) Perform steps 1. (a) through 1. (f) above.

If indicated air speed is greater than 5 knots,
the manual update will not remove 100% of
the positional error or zero out the cardinal
headings, time to go (TTG) or distance to
destination. The percentage of actual update
is a dynamic function of the computer software.

(b) If the cardinal headings, time-to-go (TTG)
and range do not decrease to 0.0, verify
that both the destination and steerpoint indicators (Dx, Sx) are set to the destination
that the update is being performed on. Repeat steps 1. (a) through 1. (f).

(1) Indicated airspeed greater than 5 knots.

(a) Ensure that the displayed Destination and
Steerpoint indicators (Dx, Sx) are both set
to the destination that the update will be
performed on.
(b) Rotate the mode select switch to the UPDT
position. The page shown in Figure 3-26
will be displayed.
(c) When the aircraft is directly over the destination point, depress the MRK key. The
page shown in Figure 3-27 will be displayed.
(d) If the pilot decides to accept the update
(ACCEPT here means to tell the INU that
the positional update will be accepted) depress line select key 7 left to accept the
update (REJECT here means to tell the
INU that the positional update will not be
accepted), depress line select key 7 right to
reject the update. In either case, the page
shown in Figure 3-27 will be redisplayed.
NOTE
The following display changes are not immediate. It will take approximately 5 seconds for the data to change.
(e) Rotate the mode select switch to NAV.
(f) Observe that the cardinal heading, timeto-go (TTG), and range decrease towards
0.0, and that present position changes to
more closely reflect the coordinates stored
in the selected destination.
(g) If the mission will continue, select a new
steerpoint and proceed.
(2) Indicated air speed is 5 knots or less.

(3) Selection of the 9UPDT9 mode on the mode
select switch deletes automatic TACAN updating during the period of the manual update.
3.23.5 Annunciations.
a. System status messages appear on line 2 and the left
side of line 6 regardless of selected page. The following is
a summary of messages that are presented and the failures/
conditions they represent.

MESSAGE

CONDITION

MSC

Mission Computer has failed.

NPU

Navigation
failed.

has

INU

INU navigation processing
has failed
Attitude may be valid.

Copilot’s Altimeter-Encoder
has failed.

TCN

TACAN has failed or is off.

PFM

Post Flight maintenance is
required.

Aircraft is within two
minutes of selected steerpoint
(flashing).

Distance to
increasing.

steerpoint

Distance to
decreasing.

steerpoint

ADC

TTG

FROM
TO
SCU

LINE

Processor

is
is

Signal Converter Unit or
ARINC BUS has failed. (See
TEST page.)

3-61

TM 1-1520-237-10

MESSAGE

CONDITION

NAVRDY
(steady)

During alignment. INU attitude data and degraded nav
performance are available.

NAVRDY
(flashing)

During alignment. Full INU
nav performance is available.

ATTD

The INU is in attitude mode
due to: 1. operator selection.
2. INU failure or 3. data bus
failure. Attitude data is valid.

LINE
U P D T

D 3

S 3

6
6

R N G

2 7 . 6

T T G

3 . 6

B R G

L
MP

1 5 2 . 7
E

1 . 0

1 . 8

6
[

DEGRD

The INU is in navigate mode
and a degraded performance
alignment, not a full performance alignment was performed.
The INU is being automatically updated by the TACAN.

]

NOTE

6
UPDT

TO SELECT THIS PAGE, SET CDU MODE
SELECT SWITCH TO UPDT. WHEN THIS
PAGE IS SELECTED, TACAN UPDATING
IS DELETED.

6
AA0397
SA

DEGUPD

Degraded mode update by
TACAN.

b. Placement of annunciation on their respective lines is
shown in Figure 3-21. NPU and PFM annunciators occupy
the same location. When a failure occurs, the annunciation
will flash to attract the pilot’s attention. Pressing the
9FACK9 key causes the annunciation to go from flashing to
steady. If an LRU recovers from the failure, its annunciator
will clear.
3.24 GYRO MAGNETIC COMPASS SET AN/ASN43.
Gyro Magnetic Compass Set AN/ASN-43 provides
heading information by reference to a free directional gyro
when operating in the FREE mode, or by being slaved to
the earth’s magnetic field when operated in the SLAVED
mode. It provides heading information to the horizontal
situation indicator. Power to operate the AN/ASN-43 is
provided from the ac essential bus through circuit breakers,
marked COMP and AUTO XFMR under the general
heading AC ESNTL BUS.
3.24.1 Compass Control C-8021/ASN-75. Control
C-8021/ASN-75 (Figure 3-29) is required to synchronize

3-62

Figure 3-26. Update Page

(electrically and mechanically align) the AN/ASN-43 to the
correct magnetic heading when used in the SLAVED mode
of operation. The synchronizing knob on the control panel
may be used as a set heading knob for operation in the
FREE mode.
3.24.2 Controls and Functions. Controls for the magnetic compass set are on the front panel of the unit. The
function panel of each control is as follows:

CONTROL

FUNCTION

Null Meter

Moves left (+) or right (v) of center
to
indicate
misalignment
(synchronization) of the AN/ASN43.

Mode Selector
(SLAVEDFREE)

Selects
either
magnetically
SLAVED or FREE gyro operation
of the AN/ASN-43.

TM 1-1520-237-10

A /

D 3

S 3
M S C N P U

R N G
T T G

2 . 1
0

B R G

1 5 2 . 7

2 2

P
AM

A C C E P T

I N U

A D C T C N

1 . 0

1 . 8

A T T D

L
MP

R E J E C T

[

]

NOTE
NOTE
TO SELECT THIS PAGE, SET CDU MODE
SELECT SWITCH TO UPDT AND PRESS
MRK KEY.

AA0398
SA

Figure 3-27. Accept/Reject Page

CONTROL
Null Control
PUSH-TO-SET

SYSTEM STATUS MESSAGES APPEAR
ON LINE 2 (2) AND THE LEFT SIDE OF
LINE 6 (6) REGARDLESS OF SELECTED
PAGE (EXCEPT NPU DATA PAGE). DATA
APPEARING ON LINES 1 (1), 3 (3), 4 (4), 5
(5), 7 (7) AND 8 (8) WILL BE WHATEVER
IS APPLICABLE TO THE PAGE SELECTED.
THE FOLLOWING IS A SUMMARY OF
MESSAGES THAT ARE PRESENTED AND
THE FAILURES / CONDITIONS
THEY REPRESENT:

Figure 3-28. System Annuciators

MODE
SELECTOR

FUNCTION
Is manually pressed and turned to
null the annunciator, thereby synchronizing (electrically and mechanically aligning) the AN/ASN43. Turns compass card of HSI for
alignment.

AA0399

NULL
CONTROL

COMPASS
+

SLAVED

FREE

PUSH TO
SET

3.24.3 Operation.
NULL METER

3.24.4 Starting Procedure.

AA0527
SA

1. Mode selector - As desired.
2. Null control - Push, and turn in direction indicated by null meter (+ or v) until annunciator is
centered. In SLAVED mode, during normal
operation, the annunciator will oscillate slightly
about the center position; however, during certain helicopter maneuvers the annunciator will
move off center.
3. HSI - check to see that HSI heading agrees with
a known magnetic heading.

Figure 3-29. Compass Control Panel C-8021/
ASN-75
3.25 ELECTRONIC
DISPLAY SYSTEM.

NAVIGATION

INSTRUMENT

The instrument display system provides displays for
navigation and command signals on a vertical situation indicator (VSI) and a horizontal situation indicator (HSI) for
pilot visual reference. The system consists of the two VSIs
and two HSIs on the instrument panel. The system has a

3-63

TM 1-1520-237-10

common command instrument system processor (CISP),
two HSI/VSI mode select panels, and one CIS mode select
panel.
3.25.1 Vertical Situation Indicator. The VSI (Figure
3-30), provides a cockpit display of the helicopter’s pitch,
roll attitude, turn rate, slip or skid, and certain navigational
information. It accepts command instrument system processor signals and displays the flight command information
needed to arrive at a predetermined point. The system also
monitors and displays warnings when selected navigation
instrument readings lack reliability. The VSI is composed
of a miniature airplane, four warning indicator flags ATT,
GS, NAV and CMD, two trim knobs ROLL and PITCH,
a bank angle scale, a bank angle index on the spheroid, a
turn rate indicator and inclinometer, pitch and roll command bars, collective position pointer, a course deviation
pointer, and a glide slope deviation pointer. Refer to Chapter 2, Section XIV for a description of the attitude indicating system, and turn and slip indicator. The gyro erect
switch (Figure 2-8) supplies a fast erect signal to the pilot
and copilot displacement gyros, thereby considerably reducing the time required for the gyros to reach full operating RPM. The pilot and copilot’s displacement gyros supply pitch and roll attitude signals to the vertical situation
indicators, automatic flight control system, and the Doppler
navigation system. Power to operate the VSI is provided
from the No. 2 ac primary bus through circuit breakers
marked VSI PLT, CPLT.
3.25.1.1 Steering Command Bars and Pointer. The
roll and pitch command bars and the collective position
pointer operate in conjunction with the command instrument system processor (CISP) and the command instrument system/mode selector (CIS MODE SEL). Selection
of HDG on the CIS MODE SEL panel provides a display
of a roll signal by the roll command bar (Figure 3-30). The
pitch command bar and the collective position pointer are
out of view, and the CMD flag is held from view. Selecting
the CIS MODE SEL switch NAV and the MODE SEL
switch VOR ILS, the roll command bar will display roll
commands from the CISP. If an ILS (LOC) frequency is
tuned in, the pitch command bar and the collective command pointer will also display CISP signals. If a VOR frequency is tuned-in, the pitch command bar and collective
position pointer will be held from view. The CMD warning
flag will be held from view, indicating that the CISP functional integrity is being monitored. Refer to Figure 3-32 for
VSI indications in other switch positions.
3.25.1.2 Command Warning Flag. The command
warning flag marked CMD is at the top left of the VSI face
(Figure 3-30). It is held from view when initial power is
applied to the CIS processor. When any CIS mode selector

3-64

switch is on, and that navigation system operating properly,
the CMD flag is not in view. During operation, if the navigation signal becomes unreliable, or is lost, the CMD flag
will become visible. On helicopters equipped with digital
CIS processor the CMD flag will not come into view when
the navigation signal becomes unreliable or lost. The NAV
flag will come into view when the navigation signal becomes unreliable even with the digital CIS.
3.25.1.3 Glide Slope Warning Flag. A glide slope
warning flag marked GS is on the right face of the indicator
(Figure 3-30). The letters GS are black on a red/white stripe
background. The warning flag will move out of view when
the ILS receivers are operating and reliable signals are received.
3.25.1.4 Navigation Warning Flag. A navigation flag
marked NAV is installed on both the VSIs and the HSIs
(Figures 3-30 and 3-31) to indicate when navigation systems are operating and reliable signals are being received.
The VSI NAV flag is marked NAV with a white background and red strips, and is on the lower left side of the
indicator. The HSI NAV flag is within the compass card
ring. Both instrument flags will retract from view whenever
a navigation receiver is on and a reliable signal is being
received.
3.25.1.5 Course Deviation Pointer. The course deviation pointer is on the VSI instrument (Figure 3-30). The
pointer works with the course bar on the HSI to provide the
pilot with an indication of the helicopter’s position with
respect to the course selected on the HSI. The scales represent right or left off course, each dot from center (on
course) is 1.25° for ILS, 5° VOR and FM. The pilot must
fly into the needle to regain on-course track.
3.25.1.6 Glide Slope Deviation Pointer. The glide
slope pointer, on the right side of the VSI (Figure 3-30), is
used with ILS. The pointer represents the glide slope position with respect to the helicopter. Each side of the on-glide
slope (center) mark are dots, each dot representing .25°
above or below the glide slope.
3.25.1.7 Controls and Indicators. Indicators of the
VSI are on the face of the instrument. The function of each
indicator is as follows:

CONTROL/
INDICATOR
Miniature
airplane/
horizon line

FUNCTION
Provides
horizon.

reference

artificial

TM 1-1520-237-10

GO−AROUND
ADVISORY
LIGHT

DECISION
HEIGHT
ADVISORY
LIGHT

ROLL
COMMAND
BAR

BANK
ANGLE
SCALE

MARKER
BEACON
ADVISORY
LIGHT

SPHERE

CMD

PITCH
COMMAND
BAR
ATT

GLIDESLOPE
DEVIATION
POINTER

CLI MB
30

BANK
ANGLE
INDEX

ARTIFICIAL
HORIZON
G
S

COLLECTIVE
POSITION
INDICATOR

NAV WARNING
FLAG (DOPPLER / GPS
VOR−LOC−FM HOMER)

NAV

MINIATURE
AIRPLANE

DI VE

PITCH
TRIM KNOB

ROLL

PITCH

ROLL TRIM KNOB
TURN
RATE
INDICATOR

INCLINOMETER

COURSE
DEVIATION
POINTER
(FM HOMER STEERING−
VOR−LOC−DOPPLER / GPS)

AA0369A
SA

Figure 3-30. Vertical Situation Indicator

CONTROL/
INDICATOR

FUNCTION

Bank angle scale

Right and left 0°, 10°, 20°, 30°, 45°,
60°, and 90° of bank.

Artificial horizon

Reference of helicopter’s attitude to
horizon.

Turn rate indicator

4-minute turn (one-needle width either side of center) 2-minute turn
(two-needle width each side of center).

Pitch and roll
command bars

Display to the pilot, control inputs
he should make to arrive at a predetermined course, or glide slope.

Collective position indicator

Display to the pilot the position of
the collective relative to where it
should be to arrive at a predetermined altitude.

CONTROL/
INDICATOR

FUNCTION

Go-around (GA) advisory light will
go on whenever the GA switch on
the pilot’s or copilot’s cyclic stick
is pressed. The light will go off
whenever the go-around mode is
ended by engaging another mode
on the CIS mode selector panel.

Decision height (DH) advisory
light will go on whenever the radar
altimeter is operating and the altitude indicator is at or below the radar altitude L (low bug) setting.

Marker beacon (MB) advisory light
will go on and the associated
marker beacon tone will be heard,
depending upon volume control setting, when the helicopter is over the
marker beacon transmitter.

Change 1

3-65

TM 1-1520-237-10

CONTROL/
INDICATOR
Glide slope
pointer

FUNCTION

CONTROL/
INDICATOR

FUNCTION

Displays to the pilot the position of
the ILS glide slope relative to the
helicopter. Pointer above center indicates helicopter is below glide
path.

COURSE set
display

Displays course to nearest degree.
Indicates same as course set
pointer.

Bearing pointer
No. 1

The pointer operates in conjunction
with Doppler/GPS or IINS.
Indicates magnetic bearing to
Doppler/GPS or IINS destination
set on FLY-TO-DEST UH or CDU
EH .

Bearing pointer
No. 2

The pointer operates in conjunction
with selected VOR or ADF
receiver. The pointer is read against
the compass card and indicates the
magnetic bearing to the VOR or
ADF station.

Course deviation
bar

This bar indicates lateral deviation
from a selected course. When the
helicopter is flying the selected
course, the course bar will be
aligned with the course set pointer
and will be centered on the fixed
aircraft symbol.

CRS knob

Course set (CRS) knob and the
course set counter operate in
conjunction with the course pointer
and allow the pilot to select any of
360 courses. Once set, the course
pointer will turn with the compass
card and will be centered on the
upper lubber line when the
helicopter is flying the selected
course.

KM indicator

Digital
distance
display
in
kilometers (KM) to destination set
on Doppler FLY TO DEST.

HDG knob

Heading set (HDG) knob operates
in conjunction with the heading
select marker, allows the pilot to
select any one of 360 headings.
Seven full turns of the knob
produces a 360° turn of the marker.

HDG warning
flag

Visible when a failure occurs in the
magnetic compass system.

To-From arrow

To-from arrow indicates that the
helicopter is flying to or away from
a selected VOR.

Course deviation
pointer

Displays to the pilot the position of
the course reference (VOR, LOC,
DPLR, FM HOME) relative to the
helicopter.

ATT warning
flag

Indicates loss of vertical gyro
power or VSI malfunction.

NAV warning
flag

Indicates loss, or unreliable signal
indication.

GS warning flag

Indicates loss, or unreliable signal
indicator.

PITCH trim
knob

Adjust artificial horizon up (climb)
from at least 4°, no more than 10°
or down (dive) from at least 8°, no
more than 20°.

ROLL trim knob

Adjust artificial horizon right or left
from at least 8° to no more than
20°.

3.25.2 Horizontal Situation Indicator. Two HSIs
(Figure 3-31) are installed on the instrument panel, one in
front of each pilot. The HSI consists of a compass card, two
bearing-to-station pointers with back-course markers, a
course bar, a KM indicator, heading set (HDG) knob and
marker, a course set (CRS) knob, a COURSE digital readout, a to-from arrow, a NAV flag, and a compass HDG
flag. The HSIs operating power is taken from the ac essential bus through a circuit breaker marked HSI PLT/CPLT.
3.25.3 Controls and Indicators. Controls of the horizontal situation indicators (Figure 3-31) are as follows:

CONTROL/
INDICATOR

FUNCTION

Compass card

The compass card is a 360° scale
that turns to display heading data
obtained from the compass control.
The helicopter headings are read at
the upper lubber line.

3-66

Change 9

TM 1-1520-237-10

LUBBER
NO. 1 BEARING
LINE
POINTER
(DOPPLER / GPS / IINS)

COURSE
SET
POINTER

DISTANCE
SHUTTER

HDG
SELECT
MARKER
COURSE
SET
DISPLAY

DOPPLER / GPS
DISTANCE
TO GO DISPLAY

H
D
G

NO. 2 BEARING
POINTER
(VOR−LF / ADF)

HDG
WARNING
FLAG

COURSE

HDG

CRS

HEADING
SET KNOB

COMPASS
CARD

COURSE
SET KNOB
NAV FLAG
(DOPPLER / GPS−
VOR−LOC)

TO−FROM
ARROW
(VOR)

COURSE
DEVIATION BAR
(DOPPLER − DOPPLER / GPS−VOR−LOC)

AA0328A
SA

Figure 3-31. Horizontal Situation Indicator

CONTROL/
INDICATOR
NAV flag

FUNCTION
The NAV flag at the top of the to
indicator, turns with the compass
card. The flag will retract from
view when a reliable navigation
signal is being applied to the instrument.

3.25.4 VSI/HSI and CIS Mode Selector Panels. The
mode select panels (Figure 3-32) are integrally lighted, instrument panel mounted controls for the VSI, HSI, and CIS.
The panels provide a means for selecting and displaying
various navigation functions. Power to operate the pilot’s
MODE SEL is taken from the No. 2 dc primary bus
through a circuit breaker, marked PILOT MODE SELECT. The copilot’s MODE SEL takes power from the
No. 1 dc primary bus through a circuit breaker, marked
CPLT MODE SELECT.

switch is set to TEST. The original indications may be restored by pressing the applicable switches.
3.25.4.1 Controls and Functions. Controls of the
mode selector panel (Figure 3-32) are as follows:

CONTROL
DPLR,
GPS

DPLR/

Directs Doppler UH , Doppler/GPS
lateral deviation and NAV flag
UH
signals to VSIs and HSIs.

VOR ILS

Directs VOR or ILS signals to
VSIs, and HSIs. Provides a signal
to NAV flag.

BACK CRS

Reverse polarity of back course
signal to provide directional display
for VSIs and HSIs. Provides a
signal to NAV flag.

FM HOME

Directs FM homing deviation and
flag signals to VSIs.

NOTE
The switches on the VSI/HSI and CIS mode
select panels may change state when the
caution/advisory panel BRT/DIM-TEST

FUNCTION

Change 1

3-67

TM 1-1520-237-10

MODE SEL
DPLR

VOR
ILS

BACK
CRS

FM
HOME

DPLR

VOR
ILS

BACK
CRS

FM
HOME

NORM
ALTER

PLT
CPLT

NORM
ALTER

ADF
VOR

TURN
RATE

CRS
HDG

VERT
GYRO

BRG
2

A
HDG

NAV

ALT

HDG

NAV

ALT

CIS MODE SEL

A
DPLR
GPS

MODE OF
OPERATIONS

CIS
MODE SELECTOR

HSI / VSI
MODE SELECTOR

NONE

ANY

MANUAL HEADING

HDG

ANY

ALTITUDE HOLD

ALT

ANY

VOR NAVIGATION

NAV

VOR

ILS NAVIGATION

NAV

ILS

ILS APPROACH

NAV

ILS

ILS BACK COURSE

NAV

BACK CRS

LEVEL OFF

NAV

VOR / ILS / BACK CRS

GO−AROUND

NAV

VOR / ILS

DOPPLER, DOPPLER / GPS

NAV

DPLR, DPLR / GPS

FM HOMING

NAV

FM HOME

AA0362_1A
SA

Figure 3-32. CIS Modes of Operation (Sheet 1 of 2)

3-68

Change 1

TM 1-1520-237-10

CYCLIC ROLL
COMMAND BAR

CYCLIC PITCH
COMMAND BAR

COLLECTIVE
POSITION INDICATOR

OFF SCALE

PROCESSED CYCLIC
ROLL COMMAND

OFF SCALE

PROCESSED COLLECTIVE
POSITION

PROCESSED CYCLIC
ROLL COMMAND

OFF SCALE

PROCESSED CYCLIC
ROLL COMMAND

PROCESSED CYCLIC
PITCH COMMAND

PROCESSED COLLECTIVE
POSITION

PROCESSED CYCLIC
ROLL COMMAND

PROCESSED CYCLIC
PITCH COMMAND

PROCESSED COLLECTIVE
POSITION

PROCESSED CYCLIC
ROLL COMMAND

OFF SCALE

PROCESSED CYCLIC
ROLL COMMAND

OFF SCALE
OR
PROCESSED CYCLIC
PITCH COMMAND

PROCESSED COLLECTIVE
POSITION

PROCESSED CYCLIC
ROLL COMMAND

PROCESSED CYCLIC
PITCH COMMAND

PROCESSED COLLECTIVE
POSITION

PROCESSED CYCLIC
ROLL COMMAND

OFF SCALE

PROCESSED CYCLIC
ROLL COMMAND

OFF SCALE

AA0362_2
SA

Figure 3-32. CIS Modes of Operation (Sheet 2 of 2)

3-69

TM 1-1520-237-10

CONTROL
TURN RATE
NORM

ALTR

CRS HDG
PLT

CPLT

BRG2
ADF

FUNCTION

HDG ON

Direct heading and roll signals to
CIS processor for steering commands that will allow pilot to maintain a selected heading.

Allows copilot’s turn rate gyro information to be displayed on pilot’s
VSI, or pilot’s gyro information to
be displayed on copilot’s VSI.

NAV ON

Gives heading commands to acquire and track a selected VOR,
ILS, DPLR, DPLR/GPS, or FM intercept, or to acquire and track glide
slope beam.

Provides for pilot’s omni-bearing
selector to be connected to navigation receiver and concurrent connection of pilot’s HSI course datum
and heading datum output to command instrument system processor.

ALT ON

Directs barometric pressure signals
and collective stick position signals
to CIS processor.

Provides pilot and copilot with his
own vertical gyro information displayed on his VSI.
Allows copilot’s vertical gyro information to be displayed on pilot’s
VSI, or pilot’s gyro information to
be displayed on copilot’s VSI.
Allows pilot or copilot to select
ADF on his No. 2 bearing pointer,
each independent of the other.

VOR

Allows pilot or copilot to select
VOR on his No. 2 bearing pointer,
each independent of the other.

CIS mode
selector

Selects one of three modes of operation to direct navigational signals to the CISP for Command Signal display.

3-70

CONTROL

Provides pilot and copilot with his
own turn rate gyro information displayed on his VSI.

Provides for copilot’s omni-bearing
selector to be connected to navigation receiver and concurrent connection of copilot’s HSI course datum and heading datum output to
command instrument system processor.

VERT GYRO
NORM

ALTR

FUNCTION

Change 5

3.25.4.2 Off Mode. The command instrument system off
mode (no switch legends lit) causes the cyclic roll, cyclic
pitch and collective command pointers on both vertical
situation indicators to be stowed out of view and the command warning flag on both VSIs to be biased out of view.
The CISP is in the off mode upon initial application of
electrical power, before the pilot selects either HDG, NAV
or ALT mode on the CIS mode selector. When NAV mode
is selected, the CISP remains in the off mode unless the
DPLR, DPLR/GPS, VOR ILS or FM HOME navigation
data has been selected on the pilot’s VSI/HSI mode selector. The CISP will return to the off mode whenever the
HDG, NAV, and ALT hold modes are disengaged, as indicated by the respective ON legends going off, or by turning off the associated navigation receiver. Separate modes
are manually disengaged by pressing the mode switch when
ON is lit.
3.25.4.3 Heading Mode. The heading mode processes
the heading error and roll attitude signals to supply a limited cyclic roll command, which, when followed, causes the
helicopter to acquire and track the heading manually selected on either pilot’s HSI. The processed signal causes
the VSI cyclic roll command bar to deflect in the direction
of the required control response; i.e., bar deflection to the
right indicates a coordinated right turn is required. When
properly followed, the command results in not more than
one overshoot in acquiring the selected heading and a
tracking error of not more than 2°. The processor gain provides 1° of roll command for each degree of heading error

TM 1-1520-237-10

up to a roll command limit of approximately 20°. The CISP
heading mode is engaged by momentarily pressing the
HDG switch on the pilot’s CIS mode selector, or as described in paragraph 3.25.4.5.
3.25.4.4 Altitude Hold Mode. The altitude hold mode
processes barometric pressure signals from the air data
transducer in addition to the collective stick position signal.
When the ALT switch on the pilot’s CIS mode selector is
pressed, the CISP provides collective command signals,
which, when properly followed, cause the helicopter to
maintain altitude to within plus or minus 50 feet. The altitude hold mode synchronizes on the engagement altitude
for vertical rates up to 200 feet per minute and provides
performance for altitude inputs between -1000 and +10,000
feet at airspeeds from 70 to 150 KIAS. It is possible to
engage the altitude hold mode, regardless of whether the
heading mode or navigation mode is engaged, except that
the CISP logic prevents manual selection of the altitude
hold mode whenever the NAV mode is engaged and an ILS
frequency is selected. This prevents the operator from selecting altitude hold mode during an instrument approach.
The altitude hold mode is manually engaged by pressing
the ALT hold switch (subject to above restriction) or automatically engaged as described in paragraph 3.25.4.7. The
altitude hold mode may be manually disengaged by pressing the ALT hold switch when the ON legend is lit. Altitude hold may be disengaged also by selecting any other
mode which takes priority (e.g., Go Around).
NOTE
ALT hold mode should be manually disabled during localizer, localizer backcourse,
VOR, and ADF approaches.
3.25.4.5 Navigation Mode. The CISP navigation mode
is engaged by pressing the NAV switch on the CIS Mode
Selector. This navigation mode causes the CISP to enter the
VOR NAV, ILS NAV, DPLR NAV, or FM HOME mode
as selected on the pilot’s VSI/HSI mode selector. The CISP
provides steering commands based on the course selected
on either the pilot’s or copilot’s HSI dependent on the mode
select CRS HDG selection of PLT or CPLT.
3.25.4.6 VOR NAV Mode. The VOR NAV mode is established by selecting the VOR/ILS switch on the VSI/HSI
mode selector and pressing the NAV switch on the CIS
mode selector. The CISP processes the heading and course
signals derived from either the pilot’s or the copilot’s HSI
in addition to the lateral deviation and lateral flag signals
applied to the pilot’s VSI. The CISP provides a limited
cyclic roll command, which, when followed, shall cause the
helicopter to acquire and track the course setting manually

selected on the HSI. Engagement of the VOR NAV when
the helicopter position is in excess of 10° to 20° from the
selected radial will cause the initial course intersection to
be made in the heading mode as described in paragraph
3.25.4.3. The CISP logic will light the CIS mode selector
HDG switch ON legend during the initial course intersection. When the helicopter is within 10° to 20° of the selected course, the CISP beam sensor will capture the VOR
lateral beam. The processor logic will turn off the HDG
switch ON legend and the final course interception, about
45°, acquisition, and tracking will be based on the VOR
lateral deviation signals. The processor causes the roll command pointer to deflect in the direction of the required control response. When properly followed, the command will
result in not more than one overshoot at a range of 10 NM
at a cruise speed of 100 6 10 knots, and not more than two
overshoots at ranges between 5 and 40 NM at speeds from
70 to 140 knots. When passing over the VOR station, the
CISP reverts to a station passage submode and remains in
this submode for 30 seconds. Cyclic roll commands during
the station passage submode will be obtained from the HSI
course datum signal. Outbound course changes may be
implemented by the HSI CRS SET knob during the station
passage submode. Course changes to a new radial, or identification of VOR intersections may be made before station
passage by setting the HSI HDG control to the present
heading and actuating the HDG switch. This will disengage
the NAV mode and allow the pilot to continue on the original radial in the heading mode. A VOR intersection fix or
selection of a new radial course may be made without affecting the CIS steering commands. Actuating the NAV
switch re-engages the VOR NAV mode to either continue
on the original VOR radial or to initiate an intercept to the
new selected radial.
3.25.4.7 ILS NAV Mode. The instrument landing system NAV mode is established by selecting the VOR/ILS
switch on the VSI/HSI mode selector, tuning a localizer
frequency on the navigation receiver and selecting the NAV
switch on the pilot’s CIS MODE SEL panel. During the
ILS NAV mode the CISP processes the following signals
in addition to those processed during the VOR NAV mode:
1. The vertical deviation and vertical flag signals, 2. the
indicated airspeed (IAS) and barometric altitude signals,
and 3. the collective stick position sensor and helicopter
pitch attitude signals. The indicated airspeed and pitch attitude signals are processed to provide a limited cyclic pitch
command, which, when properly followed, will result in
maintaining an airspeed that should not deviate more than 5
knots from the IAS existing at the time the ILS NAV mode
is engaged. The pitch command bar will deflect in the direction of the required aircraft response, i.e., an upward
deflection of the pitch bar indicates a pitch up is required.
The BAR ALT and collective stick position signals are pro-

Change 5

3-71

TM 1-1520-237-10

cessed to provide a limited collective position indication,
which, when properly followed, will cause the helicopter to
maintain the altitude existing at the time the ILS NAV
mode is engaged. The collective position indicator will deflect in the opposite direction of the required control response, i.e., an upward deflection of the collective position
indicator indicates a descent is required. The CISP will
cause the ALT hold switch ON legend to light whenever
the altitude hold mode is engaged. Actuating the ALT hold
ON switch will disengage the altitude hold mode. Desired
approach runway course must be set on the CRS window
of the HSI selected by the PLT/CPLT indication of the
CRS HDG switch. The initial course intersection and the
localizer course interception, about 45°, acquisition, and
tracking will be done as described for the VOR NAV mode
except that not more than one overshoot at a range of 10
NM at 100 6 10 KIAS, and not more than two overshoots
at ranges between 5 and 20 NM should occur for airspeeds
between 70 and 130 KIAS.
3.25.4.8 Approach Mode. The approach mode, a submode of the ILS NAV mode, will be automatically engaged when the helicopter captures the glide slope. During
the approach mode, the CISP processes the vertical deviation, GS flag, and collective stick position signals to provide a limited collective position indicator, which, when
properly followed, shall cause the helicopter to acquire and
track the glide slope path during an approach to landing.
When the glide slope is intercepted, the CISP logic disengages the altitude hold mode and causes the ON legend of
the ALT hold switch to go off. The CISP will provide a
down movement of the collective position indicator to advise the pilot of the transition from altitude hold to glide
slope tracking, and to assist in acquiring the glide slope
path. The cyclic roll commands are limited to 6 15° during
the approach submode. When properly followed, the roll
commands will result in the helicopter tracking the localizer to an approach. The collective position indicator, when
properly followed, will result in not more than one overshoot in acquiring the glidepath and have a glidepath tracking free of oscillations. The cyclic roll and collective steering performance is applicable for approach airspeed from
130 KIAS down to 50 KIAS.
3.25.4.9 BACK CRS Mode. The back course mode is a
submode of the ILS NAV mode and is engaged by concurrent ILS ON and BACK CRS ON signal from the pilot’s
HSI/VSI mode selector. The CISP monitors the localizer
lateral deviation signals to provide cyclic roll commands,
which, when properly followed, will allow the pilots to
complete back course localizer approach in the same manner as the front course ILS. The desired final approach
course should be set on the selected HSI CRS window.

3-72

Change 5

3.25.4.10 Level-Off Mode. The level-off mode will be
activated when either the VOR NAV or ILS NAV modes
are engaged, and will be deactivated by selection of another
mode or when a radar altitude valid signal is not present.
The level-off mode is not a function of a VOR or ILS CIS
approach. During ILS or VOR approaches, the barometric
altimeter must be used to determine arrival at the minimum
altitude. Radar altimeter setting shall not be used for level
off commands in the VOR NAV/ILS NAV modes because
variations in terrain cause erroneous altitude indications.
The level-off mode provides the pilots with a selectable low
altitude command. This mode is automatically engaged
when the radar altitude goes below either the pilot’s or
copilot’s radar altimeter low altitude warning bug setting,
whichever is at the higher setting. A DH legend on the VSI
and a LO light display on the radar altimeter indicator goes
on whenever the radar altitude is less than the LO bug
setting. The CISP monitors the radar altimeter and the collective stick position sensor to provide a collective pointer
command, which, when properly followed, will cause the
helicopter to maintain an altitude within 10 feet of the low
altitude setting for settings below 250 feet, and 20 feet for
settings above 250 feet. The CISP causes the ALT switch
ON legend to light and the altitude hold mode to be engaged.
3.25.4.11 Go-Around Mode. The go-around mode processes roll and pitch attitude, altitude rate, collective stick
position, and airspeed inputs in addition to internally generated airspeed and vertical speed command signals to provide cyclic roll, cyclic pitch and collective position indication. The go-around mode will engage when either pilot
presses the GA (Go Around) switch on his cyclic control
grip. When the go-around mode is engaged, the CISP immediately provides a collective position indication, which,
when followed, will result in a 500 6 50 fpm rate-of-climb
at zero bank angle. Five seconds after the GA switch is
pressed, the CISP will provide cyclic pitch bar commands,
which, when followed, will result in an 80-KIAS for the
climbout. The go-around mode is disengaged by changing
to any other mode on the pilot’s CIS mode selector.
3.25.4.12 Doppler, Doppler/GPS Mode. The Doppler, Doppler/GPS navigation mode is engaged by selecting
the DPLR, DPLR/GPS switch on the VSI/HSI mode selector and the NAV switch on the CIS mode selector. Doppler and GPS combined navigation is the default setting on
the AN/ASN-128B, but Doppler only or GPS only navigation can be selected from the DLPR/GPS CDU GPS .
During the Doppler, Doppler/GPS navigation mode, the
CISP processes Doppler, Doppler/GPS track angle error
and the Doppler, Doppler/GPS NAV flag signals in addition to the roll angle input from the attitude gyro. The CISP

TM 1-1520-237-10

provides cyclic roll bar commands, which, when followed,
result in a straight line, wind-corrected, flight over distances

greater than 0.2 kilometer from the destination. The course

Change 5

3-72.1/(3-72.2 Blank)

TM 1-1520-237-10

deviation bar and course deviation pointer provide a visual
display of where the initial course lies in relationship to the
helicopter’s position. The initial course is the course the
Doppler, Doppler/GPS computes from the helicopter’s position to the destination at the time the fly to destination
thumbwheel is rotated (or entered from the keyboard). The
VSI and HSI course sensitivity is 61000 meters when farther than 12 km from the fly-to destination. Course sensitivity gradually scales down from 61000 meters at 12 km
to 6200 meters at 2 km and less from the fly-to destination.
To achieve a pictorially correct view of the course, rotate
the course knob to the head of the No. 1 needle when the
fly to destination thumbwheel is rotated (or entered from
the keyboard). The DPLR, DPLR/GPS NAV logic detects
the condition of station passover, and automatically
switches to heading mode. The switch to heading mode will
be indicated by the HDG switch ON legend being turned
on, and the NAV switch ON legend being turned off. The
Doppler, Doppler/GPS navigation mode will not automatically re-engage, but will require manual re-engagement of
the NAV switch on the CIS mode selector.
3.25.4.13 FM HOME Mode. The FM homing (Figure
3-32) is engaged by selecting the FM HOME switch on the
pilot’s VSI/HSI mode selector and the NAV switch on the
pilot’s CIS mode selector. Selecting FM homing on the
VSI/HSI mode selector directs FM homing signals only to
the VSI. Other NAV modes will be retained on the HSI if
previously selected. During the FM HOME mode, the
CISP processes the lateral deviation and flag signals displayed on the pilot’s VSI in addition to the roll angle input
from the attitude gyro. The CISP filters and dampens the
FM homing deviation signals and provides cyclic roll commands to aid the pilot in homing on a radio station selected
on the No. 1 VHF-FM communications receiver. When
properly followed, the roll commands result in not more
than two overshoot heading changes before maintaining a
tracking error not to go over 3°. The CISP will revert to the
heading mode whenever the lateral deviation rate is over
1.5°/ sec for a period of over 1 second. The CISP will cause
the CIS mode selector HDG switch ON legend to light, and
remain in the heading mode until the FM mode or some
other mode is manually selected. Concurrent VOR and FM
or concurrent DPLR and FM mode inputs will be considered an FM mode input to the CISP.
3.25.4.14 TURN RATE Select. The turn rate gyro selection provides each pilot the option of having his VSI
display his own turn rate gyro signal (NORM operation) or
of having the other pilot’s turn rate gyro signal displayed
(ALTR operation). The turn rate gyro selection is independent of the navigation modes selected by the top row of
switches and is independent of which turn rate gyro the
other pilot has selected. The NORM selection connects

each pilot’s VSI to his own turn rate gyro. The selection of
NORM or ALTR operation is indicated by lighting the
respective legend on the TURN RATE selector switch.
The lamp power to the indicator legends is controlled
through a relay so that the NORM legend is lit in case the
mode selector logic or lamp drivers fail. Sequential operation of the TURN RATE switch alternates the rate gyro
connected to the VSI.
3.25.4.15 CRS HDG Select. The CRS HDG switch on
the mode selector provides for either the pilot’s or the copilot’s course selector (CRS) to be connected to the navigation receiver, and for concurrent connection of the same
pilot’s HSI course and heading information to the command instrument system processor. The CRS resolver is
normally connected to the pilot’s HSI until selected by the
copilot on his mode selector. CRS HDG control is transferred by pressing the CRS HDG switch. The pilot having
the CRS HDG control is indicated by lighting of either the
PLT or the CPLT legend on each mode selector. When
power is first applied to the mode selector, the pilot’s position is automatically selected. The CRS HDG selection is
independent of the navigation modes selected by the top
row of switches.
3.25.4.16 VERT GYRO Select. The vertical gyro selection provides each pilot the option of having his VSI
display his own vertical gyro attitude (NORM operation),
or of having the other pilot’s vertical gyro attitude displayed (ALTR operation). The vertical gyro selection is
independent of the navigation modes selected by the top
row of switches and is independent of which vertical gyro
the other pilot has selected. Each pilot’s VSI is normally
connected to his own vertical gyro. The selection of
NORM or ALTR operation is indicated by lighting the
respective legend on the VERT GYRO selector switch.
The lamp power to the indicator legends is controlled
through a relay so that the NORM legend is lit in case the
mode selector logic or lamp drivers fail. Sequential operation of the VERT GYRO switch alternates the vertical
gyro connected to the VSI.
3.25.4.17 No. 2 Bearing Select. The HSI number 2
bearing pointer selection allows the option of either the
LF/ADF bearing or the VOR bearing to a selected station.
The ADF/VOR selection is independent of the navigation
modes selected by the top row of switches, and either pilot
selects ADF or VOR, independent of the other pilot’s selection. The number 2 bearing pointer is normally connected to the LF/ADF bearing output. The selection of either ADF or VOR bearing is indicated by lighting of the
respective legend on the selector switch. The lamp power to
the indicator legends is controlled through a relay, so that
the ADF legend is lit in case the mode selector logic or

Change 10

3-73

TM 1-1520-237-10

lamp drivers fail. Sequential operation of the ADF/VOR
switch alternates the bearing source connected to the No. 2
bearing pointer between ADF or VOR.
3.25.5 Operation.
a. Heading Hold.
(1) CIS MODE SEL switch - HDG.
(2) HDG set knob on HSI - Set as desired.
(3) Selected heading is achieved by banking
helicopter, to center roll command bar.

(3) CIS MODE SEL switch - NAV.
(4) At two dots localizer deviation on HSI, follow roll command bar to intercept localizer.
(5) As glide slope deviation pointer centers,
follow collective position indications for
glide slope tracking.
(6) At decision height, press GA switch for
go-around mode if breakout has not occurred.
d. Back Course Localizer Approach.

b. VOR Course Intercept.
(1) Frequency - Set.
(1) Frequency - Set.
(2) HSI CRS set knob - Set to desired course.
(3) CIS MODE SEL switch - NAV.
(4) Follow roll command bar to initially follow intercept heading and then follow command bar to intercept VOR course.
c. ILS Approach.
(1) Frequency - Set.
(2) HSI CRS set knob - Set to desired course.

3-74

Change 10

(2) LO altitude bug - SET to missed approach
point HAT.
(3) HSI CRS set knob - Set to inbound back
course.
(4) CIS MODE SEL switch - NAV.
(5) MODE SEL switch - BACK CRS.
(6) Fly same as front course (paragraph
3.25.5c(4)). Turn off MODE SEL ALT
legend to stow collective position indicator
before making manual descent on back
course approach.

TM 1-1520-237-10

Section IV TRANSPONDER AND RADAR
3.26 TRANSPONDER AN/APX-100(V)1 (IFF).
The transponder set (Figure 3-33) provides automatic
radar identification of the helicopter to all suitably equipped
challenging aircraft and surface or ground facilities within
the operating range of the system. AN/APX-100(V) receives, decodes, and responds to the characteristic interrogations of operational modes 1, 2, 3/A, C, and 4. Specially
coded identification of position (IP) and emergency signals
may be transmitted to interrogating stations when conditions warrant. The transceiver can be operated in any one of
four master modes, each of which may be selected by the
operator at the control panel. Five independent coding
modes are available to the operator. The first three modes
may be used independently or in combination. Mode 1 provides 32 possible code combinations, any one of which
may be selected in flight. Mode 2 provides 4096 possible
code combinations, but only one is available and is normally preset before takeoff. Mode 3/A provides 4096 possible codes, any one of which may be selected in flight.
Mode C will indicate pressure altitude of the helicopter
when interrogated. Mode C is only available if both mode
3/A and mode C switches are placed to the ON position.
Mode 4 is the secure mode of cooperative combat identification, IFF operational codes are installed, the current period’s code and either the previous or the next period’s
code. Power to operate the IFF system is provided from the
No. 1 dc primary bus through a circuit breaker marked IFF.
Refer to TM 11-5895-1199-12 and 11-5895-1037-12.

switch is in the TOP position and the stronger signal was
received from the bottom antenna, no rf reply will be transmitted. If the ANT switch is in the BOT position and the
stronger signal was received from the top antenna, no rf
reply will be transmitted. Therefore the ANT switch must
be in the DIV position to ensure the IFF will reply to all
valid interrogations.
3.26.2 Controls and Functions. All operating and
mode code select switches for transceiver operation are on
Control Panel RT-1296/APX-100(V) (Figure 3-33).
CONTROL/
INDICATOR
TEST GO

Indicates successful BIT.

TEST/MON NO
GO

Indicates unit malfunction.

ANT-DIV switch

Allows the pilot to select the TOP
(upper antenna), BOT (bottom antenna), or DIV (diversity, both antennas) of the aircraft.
NOTE

The ANT-DIV switch shall be
placed in the DIV position at all
times.
MASTER/OFF/
STBY/NORM/
EMER

3.26.1 Antenna.

CAUTION

The transponder will ignore (and not respond to) interrogations received from the
ground if the ANT switch is in the TOP
position and will ignore interrogations received from above if the ANT switch is in
the BOT position.
Flush-mounted antennas are installed on the top fairing
between engine exhaust ports, (Figure 3-1) and under the
transition section behind the UHF-AM antenna. They receive signals of interrogating stations and transmit reply
signals. The AN/APX-100(V) is a diversity transponder,
functioning to receive the rf interrogation from two antennas and transmit the reply via the antenna from which the
stronger interrogation signal was received. If the ANT

FUNCTION

Selects

operating

condition.

NOTE

Emergency reply provisions. This
mode of operation is possible when
the MASTER switch on the IFF
control panel is placed in the
EMER position and the system is
interrogated. Emergency operation
results in four short dashes on the
interrogating radar indicator, which
indicates an aircraft in distress, and
singles out the aircraft in emergency condition within the group of
aircraft. (The MASTER switch
must be in NORM, then lifted and
turned to EMER, therefore preventing the switch from accidentally being in EMER.) The emergency reply consists of a code 7700
in mode 3/A.

Change 10

3-75

TM 1-1520-237-10

MODE 3A
FUNCTION
SWITCH
TEST / MON
NOGO
INDICATOR

MODE 2
FUNCTION
SWITCH

ANTENNA
SELECTOR
SWITCH

MODE 4
TEST−ON−OUT
SWITCH

T
TO T

TOP
A
N
T

BOT

TEST

M−3/A

RAD
TEST

M−C

O
N

MASTER
D
I
V

N ORM S T

M−2

N
O
G
O

STATUS
INDICATOR
ALTITUDE
DIGITIZER

OFF
BY

O
N

STATUS
INDICATOR
EXTERNAL
COMPUTER

STATUS

OUT
ALT

CODE

RO B
ZE

P R ES

L
I
G
H
T

O
N

OUT

AUDIO

IDENT

TEST

TO T

MIC

OUT

MODE 1

ANT

A HO

I
F
F

KIT

MODE 4
LD

MODE 4
CODE
HOLD−A−B−ZERO
SWITCH

O
U
T

STATUS
INDICATOR
ANTENNAS
MODE C
FUNCTION
SWITCH

MODE 3 / A

MASTER
CONTROL
SWITCH

TO T

M−1

P R ES

MODE 1
FUNCTION
SWITCH

TEST / MON

TEST
G
O

RAD
TEST−OUT
SWITCH

TEST GO
INDICATOR

IDENTIFI−
CATION
POSITION
(IP)
MODE 4
REPLY
INDICATOR

MODE 2
CODE
SELECTOR
BUTTON

MODE 1
CODE
SELECTOR
BUTTON

MODE 3A
CODE
SELECTOR
BUTTON

MODE 2
CODE
SELECTOR
BUTTON

MODE 3A
CODE
SELECTOR
BUTTON

MODE 4
AUDIO−LIGHT−OUT
SWITCH
AA0363A
SA

Figure 3-33. Control Panel RT-1296/APX-100(V)

3-76

Change 1

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

M-1, M-2,
M-3/A, M-C
switches

The four, three position switches on
the IFF control panel will enable or
disable the system for modes 1, 2,
3/A, or C operation. Mode 1, mode
2, mode 3/A, or mode C replies are
possible only when their respective
switches are placed in the ON positions. Mode C is available only if
both mode 3/A and mode C are
placed in the ON position. Mode 1
switches permit selection of a desired code from 00 through 73.
Mode 2 and mode 3/A switches
permit selection of a desired code
from 0000 through 7777. The OUT
position if each switch prevents a
reply to the respective mode interrogations. The TEST position of
each switch tests the respective
mode operation.

RAD TEST/
OUT

RAD TEST
OUT

The RAD switch is used to allow
the RT to reply to external test interrogation when held in the RAD
position.
Allows receiver transmitter to reply
to external test interrogations.

CONTROL/
INDICATOR
MODE 4 CODE
selector

When the IFF mode 4 computer is
installed, mode 4 interrogations bypass the decoder in the RT and go
directly to the crypto computer. In
the crypto computer the mode 4 interrogation signal is decoded and
applied to the mode 4 recognition
circuit. When a mode 4 complete
concurrence exists, the mode 4 recognition circuit generates a signal
to the mode 4 computer which in
turn generates a mode 4 reply. The
REPLY light on the IFF control
unit comes on to indicate a mode 4
reply is being transmitted.

ZERO

Zeroize code setting in computer.

Selects mode 4 code setting for previous, present, or next period, depending on which crypto period applies.

HOLD

Retains mode 4 code setting when
power is removed from transponder.

Disables the RAD TEST features
of the transponder.

STATUS ALT

Indicates that BIT or MON failure
is due to altitude digitizer.

STATUS KIT

Indicates that BIT or MON failure
is due to external computer.

STATUS ANT

Indicates that BIT or MON failure
is due to cables or antenna.

FUNCTION

MODE 4 TEST/
ON/OUT
ON

Allows system to reply to mode 4
interrogations.

OUT

Prevents reply to mode 4 interrogations.

TEST

Provides self test for mode 4.

MODE 4
AUDIO/LIGHT/
OUT
AUDIO

Enables aural and REPLY light
monitoring of valid mode 4 interrogations and replies. (Preferred position)

LIGHT

Enables only REPLY light monitoring of valid mode 4 interrogations and replies.

Change 1

3-77

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

WARNING
Placing the switch in the OUT
position will disable mode 4 REPLY monitoring and IFF caution
light.
OUT

Disables aural, REPLY light, and
caution light monitoring of valid
mode 4 interrogations and replies.

MODE 4 REPLY

Indicates that a mode 4 reply is
transmitted.

IDENT/
OUT/MIC

The IDENT/OUT/MIC switch is
spring loaded to the OUT position.
If IDENT operation is desired, the
switch must be moved to the
IDENT position momentarily. The
IDENT pulse trains will be transmitted for approximately 30 seconds. The MIC position is not connected in this installation.

MODE 1 selector
buttons

Selects mode 1 reply code to be
transmitted.

MODE 2 selector
buttons

Selects four digit mode 2 reply code
to be transmitted. (Located on the
control panel or on the remote RT.)

MODE 3/A selector buttons

Selects four digit mode 3/A reply
code to be transmitted.

3.26.3 Operation.
3.26.3.1 Starting Procedure.

If the MODE 2 code has not been set previously, loosen
two screws which hold MODE 2 numeral cover, and slide
this cover upward to expose numerals of MODE 2 code
switches (Figure 3-33). Set these switches to code assigned
to helicopter. Slide numeral cover down and tighten screws.
1. MASTER switch - STBY. NO-GO light
should be on.
2. Allow 2 minutes for warmup.
3. MODES 1 and 3A CODE selector buttons Press and release until desired code shows.
4. TEST, TEST/MON, and REPLY indicators
-PRESS-TO-TEST. If MODE 1 is to be used,
check as follows:
5. ANT switch - DIV.
6. MASTER switch - NORM.
7. M-1 switch - Hold at TEST, observe that only
TEST GO indicator is on.
8. M-1 switch - Return to ON. If modes 2, 3A or
M-C are to be used, check as follows:
9. M-2, M-3/A and M-C switches - Repeat steps
7. and 8.
NOTE
Do not make any checks near a radar site or
with MASTER control switch in EMER,
nor with M-3/A codes 7500, 7600 or 7700,
without first obtaining authorization from the
interrogating station(s) within range of the
transponder.
The following steps can be done only with
KIT/1A computer transponder installed.
10. MODE 4 CODE switch - A.

CAUTION

a. Set assigned test code in the KIT/1A computer transponder.
When flying in a combat situation near
friendly radar sites or in the vicinity of
friendly fighter aircraft, the MODE 4
monitor switch must be in either the AUDIO or LIGHT position. This will enable
the pilot to observe that the IFF is periodically responding to expected MODE 4
interrogations.
3-78

Change 8

b. AUDIO-ON-OUT switch - OUT.
c. MODE 4 TEST-ON-OUT switch - Place
to TEST and hold, then release.
d. TEST GO light - ON, MODE 4 REPLY
light off, KIT STATUS light off.

TM 1-1520-237-10

11. When possible, request cooperation from interrogating station to activate radar TEST mode.
a. Verify from interrogating station that
MODE TEST reply was received.
b. RAD TEST switch - RAD TEST and
hold.
c. Verify from interrogating station that
TEST MODE reply was received.
3.26.3.2 Normal Procedures. Completion of the starting procedure leaves the AN/APX-100(V) in operation. The
following steps may be required, depending upon mission.
1. MODE 4 CODE selector switch - A or B as
required.
a. If code retention is desired, momentarily
place the MODE 4 CODE selector switch
to HOLD prior to turning the MASTER
switch OFF.
b. If code retention in external computer is
not desired during transponder off mode,
place MODE 4 CODE selector switch to
ZERO to dump external computer code
setting.
2. Mode M-1, M-2, M-3/A, M-C, or MODE 4
switches - Select desired mode.
3. Identification of position (I/P) switch - IDENT,
when required, to transmit identification of position pulses.
3.26.3.3 Emergency Operation.

3.27 TRANSPONDER COMPUTER KIT-1A/TSEC.
The transponder computer in the nose section of the helicopter operates in conjunction with mode 4. A caution
light on the caution panel, marked IFF, will go on when a
malfunction occurs in mode 4 or the computer that will
prevent a reply when interrogated. Mode 4 operation is
selected by placing the MODE 4 switch ON, provided the
MASTER switch is at NORM. Placing the MODE 4
switch to OUT disables mode 4. MODE 4 CODE switch
is placarded ZERO, B, A, and HOLD. The switch must be
lifted over a detent to switch to ZERO. It is spring-loaded
to return from HOLD to the A position. Position A selects
the mode 4 code for the previous, present, or next period
depending on which crypto period applies and position B
selects the mode 4 code for previous, present, or next period depending on which crypto period applies. Both codes
are mechanically inserted by a code-changing key. The
codes are mechanically held in the transponder computer,
regardless of the position of the MASTER switch or the
status of helicopter power, until the first time the helicopter
becomes airborne. Thereafter, the mode 4 codes will automatically zeroize any time the MASTER switch or helicopter power is turned off. The code setting can be mechanically retained. With weight on the landing gear, turn
the MODE 4 CODE switch to HOLD (only momentary
actuation is required) and release. Mode 4 codes can be
zeroized any time the helicopter power is on and the MASTER switch is not in OFF, by turning the CODE switch to
ZERO. Power to operate the transponder computer is provided automatically when the AN/APX-100(V) is on. The
transponder computer KIT-1A/TSEC operation is classified.
3.28 CRYPTOGRAPHIC COMPUTER KIT-1C.
The cryptographic computer uses electronic key loading.
Key loading is accomplished by use of the KYK-13 Electronic Transfer Device per TM 11-5810-389-13&P. The
Cryptographic Computer Kit-1C operation is classified.

NOTE
3.29 RADAR ALTIMETER SET AN/APN-209(V).
MASTER control switch must be lifted before it can be switched to NORM or EMER.
During a helicopter emergency or distress condition the
AN/APX-100(V) may be used to transmit specially coded
emergency signals on mode 1, 2, 3/A and 4 to all interrogating stations. Those emergency signals will be transmitted as long as the MASTER control switch on the control
panel remains in EMER and the helicopter is interrogated.
MASTER control switch - EMER.
3.26.4 Stopping Procedure. MASTER switch - OFF.

The radar altimeter set (Figure 3-34) provides instantaneous indication of actual terrain clearance height. Altitude,
in feet, is displayed on two radar altimeter indicators on the
instrument panel in front of the pilot and copilot. The radar
altimeter indicators each contain a pointer that indicates
altitude on a linear scale from 0 to 200 feet (10 feet per
unit) and a second-linear scale from 200 to 1500 feet (100
feet per unit). An on/OFF/LO altitude bug set knob, on the
lower left corner of each indicator, combines functions to
serve as a low level warning bug set control, and an on/
OFF power switch. The system is turned on by turning the

Change 8

3-78.1

TM 1-1520-237-10

3.29.1 Antennas. Two identical radar altimeter antennas (Figure 3-1) are on the cockpit section under the avionics compartment. One is for the transmitter and the other
is for the receiver. The antennas are flush-mounted in the
fuselage on the bottom of the helicopter.
3.29.2 Controls or Indicator Function. Control of the
radar altimeter set is provided by the LOW SET OFF knob
on the front of the height indicator. The knob, marked HI
SET, also controls the PUSH TO TEST (Figure 3-34).

LO
SET
BUG
H

LO control knob, marked SET, of either indicator, clockwise from OFF. Continued clockwise turning of the control
knob will permit either pilot to select any desired lowaltitude limit, as indicated by the LO altitude bug. Whenever the altitude pointer exceeds low-altitude set limit, the
LO altitude warning light will go on. Pressing the PUSHTO-TEST HI SET control provides a testing feature of the
system at any time and altitude. When the PUSH-TOTEST control knob is pressed, a reading between 900 feet
and 1100 feet on the indicator, and a reading between 900
and 1100 feet on the digital display, and the OFF flag
removed from view, indicates satisfactory system operation. Releasing the PUSH-TO-TEST SET control knob
restores the system to normal operation. A low-altitude
warning light, on the center left of the indicator, will light
to show the word LO any time the helicopter is at or below
the altitude limit selected by the low altitude bug. Each
pilot may individually select a low-altitude limit and only
his LO light will go on when the low-altitude is reached or
exceeded. Loss of system power will be indicated by the
indicator pointer moving behind the dial mask and the OFF
flag appearing in the center of the instrument. If the system
should become unreliable, the flag will appear and the indicator pointer will go behind the dial mask, to prevent the
pilot from obtaining erroneous readings. Flight operations
above 1600 feet do not require that the system be turned
off. The pointer will go behind the dial mask but the transmitter will be operating. Power to operate the AN/APN-209
is supplied from the No. 1 dc primary through circuit
breakers, marked RDR ALTM.

HI
WARNING
LIGHT

ABS

LO
WARNING
LIGHT

FT X 100

ALTITUDE
POINTER

HI
SET
BUG

ALT

SET

DIAL
MASK

SET
PUSH
TO TEST

OFF

DIGITAL
READOUT

FEET

SYSTEM
OFF FLAG

AA0528
SA

Figure 3-34. Radar Altimeter Set AN/APN-209(V)

CONTROL/
INDICATOR

FUNCTION OR INDICATION

HI SET knob

Pushing knob actuates built-in test
system to self-test altimeter.

Altitude pointer

Provides an analog indication of
absolute altitude from zero to 1500
feet.

Digital readout

Gives a direct-reading four digit
indication of absolute altitude from
zero to 1500 feet.

LO warning light

Lights whenever dial pointer goes
below L altitude bug setting.

HI warning light

Lights whenever dial pointer goes
above H altitude bug setting.

OFF flag

Moves into view whenever
altimeter loses track while power is
applied.

3.29.3 Operation.
CONTROL/
INDICATOR

FUNCTION OR INDICATION

LO SET knob

Power
control
counterclockwise
to
clockwise to on.

turned
OFF,

a. Starting Procedure.
(1) LO SET knob - On.
(2) L bug - Set to 80 feet.

L bug

Sets altitude trip point of LO
warning light.

(3) H bug - Set to 800 feet.

H bug

Sets altitude trip point of HI
warning light.

(4) Indicator pointer - Behind mask above
1500 feet.

3-78.2

Change 8

TM 1-1520-237-10

b. Track Operation. After about 2 minutes of warmup, the altimeter will go into track mode with
these indications:
(1) OFF flag - Not in view.

3.29.4 Stopping Procedure.
LO SET knob - OFF.
3.30 MISSION EQUIPMENT INTERFACE.

(2) Altitude pointer - 0 6 5 feet.
CAUTION

(3) Digital readout - 0 to +3 feet.
(4) LO warning light - Will light.
(5) HI warning light - Will be off.
c. HI SET knob - Press and hold. The altimeter
will indicate a track condition as follows:
(1) OFF flag - Not in view.
(2) Altitude pointer - 1000 6 100 feet.
(3) Digital readout - 1000 6 100 feet.
(4) LO warning light - Will be off.
(5) HI warning light - Will light.
(6) HI SET knob - Release. The altimeter will
return to indications in step b. Track Operation.

The ECM antenna can be extended with
the helicopter on the ground if the radar
altimeter is turned off or removed from
the installation, or the L (LO set) indicator is set below the radar altimeter indication.
Two signals are provided by the radar altimeter to the
AN/ALQ-151(V)2 mission equipment. RADAR ALTIMETER ON indicates the altimeter is installed and has power
applied. If this signal is not present, and the ECM antenna
is not fully retracted, a signal is generated to light the ANTENNA EXTENDED capsule on the CAUTION/
ADVISORY panel. The other signal, RADAR ALTITUDE LOW, is sent to the mission equipment when the
aircraft altitude drops below the LO bug setting of the radar
altimeter. The signal initiates automatic retraction of the
ECM antenna, lights the ANTENNA EXTENDED capsule
until the antenna is fully retracted, and disables the ECM
ANTENNA switch.

Change 8

3-79/(3-80 Blank)

TM 1-1520-237-10

CHAPTER 4
MISSION EQUIPMENT
Section I MISSION AVIONICS
4.1 TROOP COMMANDER’S ANTENNA.
The troop commander’s antenna (Figure 3-1), on the upper trailing edge of the tail rotor pylon, provides for use of
a VHF/FM mobile/man pack radio, such as the AN/PRC-25
or AN/PRC-77, from the cabin area. The antenna gives the
troop commander the capability of liaison, command, and
control of ground elements. A coaxial cable, coiled in the
cabin ceiling near the left cabin door, is for connecting the
antenna to the radio set.
4.2 CREW CALL SWITCH/INDICATOR.

The CREW CALL switch/indicators are on the instrument panel (Figure 4-1) and in the cabin at the DF and
ECM consoles. The switches are used to provide signals
between crew members to indicate communication is desired, and establishing ICS circuits between cockpit and
cabin. When the pilot/copilot’s CREW CALL switch is
pressed in, it lights steady. This allows only one-way communication, from pilot/ copilot to mission equipment operator(s). All stations desiring to communicate must then place
their respective intercom switches to ICS. To establish twoor three-way communications, the flashing switches must
be pressed in. The pilot’s ICS audio overrides all other
mission equipment operator’s audio. To establish communication from mission equipment operator(s) to pilot/
copilot, the DF and/or ECM operator must press in their
respective CREW CALL switch. The DF and/or ECM operator(s) CREW CALL switch(es) will light steady. The
pilot/copilot’s CREW CALL light flashes. When the pilot/
copilot’s CREW CALL switch is pressed in, the switch
lights steady, and communications can then be established.
In establishing communications, the first CREW CALL
switch pressed will light steady, all others will flash until
pressed in. To terminate two-way communication, the pilot/
copilot and mission equipment operator(s) must press the
respective CREW CALL switch(es), causing all indicators
to go off. In terminating communications, CREW CALL
switches pressed in must be pressed to release. Power to

operate the CREW CALL system is provided from the No.
1 dc primary bus through a circuit breaker marked
LIGHTS ADVSY.
4.3 CHAFF AND FLARE DISPENSER M130.
4.3.1 Chaff Dispenser M130. The general purpose dispenser M130 (Figure 4-1) consists of a single system (dispenser assembly, payload module assembly, electronics
module and dispenser control panels) and a CHAFF DISPENSE control button (on the lower console) designed to
dispense decoy chaff, M-1 (refer to TM 9-1095-206-13&P).
The system provides effective survival countermeasures
against radar guided weapon systems threats. The dispenser
system, M130, has the capability of dispensing 30 chaff.
Power to operate the chaff dispenser system is provided
from the No. 1 dc primary bus through a circuit breaker,
marked CHAFF DISP.
4.3.2 Flare Dispenser M130. EH The general purpose
dispenser (Figure 4-1) consists of a single system dispenser
assembly, payload module assembly, electronics module,
and dispense control button (on the instrument panel), designed to dispense decoy flares M206. The system provides
effective survival countermeasure against infrared sensing
missile threats. The dispenser system has the capability of
dispensing 30 flares. Power to operate the flare dispenser
system is provided from the No. 1 dc primary bus through
a circuit breaker marked CHAFF DISP.
4.3.3 Controls and Function. The dispenser control
panel (Figure 4-1) contains all necessary controls to operate
the dispenser system from the cockpit. The control panel is
on the lower console.
CONTROL/
INDICATOR
CHAFF counter

FUNCTION
Shows the number of chaff cartridges
remaining in payload module.

4-1

TM 1-1520-237-10

INFRARED
COUNTERMEASURE
TRANSMITTER

B
A

EH
ONLY

RADAR SIGNAL
DETECTOR
INDICATOR

BDHI

0 0 0 0

33
30

12
21

A
FLARE
CREW
CALL

RETRACT
OFF
EXTEND

FLARE
RELEASE
BUTTON

C
DISPENSER
ASSEMBLY

PAYLOAD
MODULE

ECM
ANTENNA
SWITCH

SWITCH / LIGHT

INSTRUMENT PANEL

B
ELECTRONIC
MODULE

CHAFF / FLARE
DISPENSER
CONTROL PANEL

FLARE
0

DISP
CONT

ARM

CHAFF

ARM

SAFETY PIN AND
WARNING STREAMER

R
I
P
P
L
E

RADAR SIGNAL
DETECTOR
CONTROL PANEL
PWR

MAN

F
I
R
E
SAFE
SELF

ON
OFF

PGRM

DSCRM
ON

TEST

OFF

AUDIO

COVER

SELECT
SWITCH
C (CHAFF)
OR F (FLARE)

CHAFF DISPENSE
SWITCH
CHAFF
DISPENSE

LOWER CONSOLE
(TYPICAL)

FLARE/CHAFF DISPENSERS

Figure 4-1. Mission Kits

4-2

AA0372A
SA

TM 1-1520-237-10

FUNCTION

CONTROL/
INDICATOR

FUNCTION

Chaff counter setting
knob

Adjusts counter to correspond to
number of chaff cartridges remaining in payload module.

DISP CONT
RIPPLE FIRE

Release (jettison) all flares from
payload module without pressing
FLARE switch for each flare.

PUSH-RESET

When pushed, resets chaff counter
to 9009.

FLARE switch
(Instrument panel)

Fires one flare from payload module each time switch is pressed.

ARM indicator
light

Indicates that arming switch is at
ARM, safety flag pin is removed,
and payload module is armed.

CONTROL/
INDICATOR

ARM-SAFE switch
ARM

Applies electrical power through
safety flag switch to CHAFF DISPENSE button, and flare firing
switch. Flare firing system is not
used in this installation.

SAFE

Removes power from dispenser
system.

FLARE counter

Not used in this installation.

Flare counter setting
knob UH

Not used in this installation.

DISP CONT

Not used in this installation.

Mode selector

Selects type of chaff release operation.

MAN

Dispenses one chaff cartridge each
time dispense button is pressed.

PGRM

Dispenses chaff according to predetermined burst/salvo and number of
salvos automatically.

CHAFF DISPENSE

Ejects chaff cartridges from payload module.

4.3.4 Controls and Function.

4.3.5 Dispenser Assembly. The dispenser assembly
(Figure 4-1) contains the breech assembly, C-F selector
switch for either chaff or flares, a reset switch, and a housing containing the sequencer assembly. The sequencer assembly receives power through the firing switches circuit
and furnishes pulses to each of the 30 contacts of the breech
assembly, in sequential order 1 through 30, thus firing each
of the impulse cartridges.
4.3.6 Payload Module Assembly. The payload module assembly (Figure 4-1) consists of the payload module
and retaining plate assembly. The payload module has 30
chambers which will accept chaff. The chaff cartridges are
loaded through the studded end of the module, one per
chamber, and are held in place by the retaining assembly.
The payload module assembly is assembled to the dispenser
assembly.
4.3.7 Electronic Module Assembly (EM). The EM
(Figure 4-1) contains a programmer and a cable assembly
which includes a 28-volt supply receptacle and a safety
switch, actuated by inserting the safety pin with streamer
assembly. The programmer consists of a programming circuit which allows the setting of chaff burst number, chaff
burst interval, chaff salvo number, and chaff salvo interval.
4.3.8 Electronics Module Controls. Controls on the
electronic module are used to program the chaff dispenser
for predetermined release of chaff cartridges. Controls on
the electronic module are as follows (refer to TM 9-1095206-13&P):

CONTROL
CONTROL/
INDICATOR

FUNCTION

FLARE counter

Indicates the number of flare cartridges remaining in payload module.

Flare counter set
knob

Adjusts counter to correspond to
number of flare cartridges in payload module.

SAFETY PIN

Safety switch to accept the safety
pin with streamer, placing the dispenser in a safe condition when the
helicopter is on the ground.

SALVO COUNT

Programs the number of salvos; 1,
2, 4, 8 or C (Continuous).

4-3

TM 1-1520-237-10

CONTROL
SALVO
VAL

INTER-

BURST COUNT

FUNCTION
Programs the time in seconds between salvos; 1, 2, 3, 4, 5, 8 or R
(Random 2, 5, 3, 4, 3).

the system is activated again by the FLARE
switch.
4.3.12 Stopping Procedure.
ARM SAFE switch - SAFE.

Programs the number of burst; 1, 2,
3, 4, 6 or 8.

4.4 DELETED.

Programs the time in seconds for
burst intervals; 0.1, 0.2, 0.3 or 0.4.

4.5 BEARING, DISTANCE, HEADING INDICATOR
(BDHI). EH

4.3.9 Safety Procedures. The safety pin shall be installed in the safety switch when the helicopter is parked.
Safety pin is removed immediately before takeoff.

The BDHI (Figure 4-1) at the center of the instrument
panel consists of three indicators. The position of the indicator allows easy viewing by both pilot and copilot. The
functions of the indicators are as follows:

BURST
VAL

INTER-

4.3.10 Operation.
1. Counter(s) - Set for number of cartridges in
payload module(s).
2. Mode switch - MAN.

a. Compass Rose - displays the magnetic heading of the
helicopter.
b. Bearing Pointer - displays bearing to the signal received from an airborne or ground emitter/transmitter. The
DF operator selects signal to be displayed.

NOTE
Mode switch should always be at MAN
when the ARM-SAFE switch is moved to
ARM to prevent inadvertent salvo of chaff.
3. ARM SAFE switch - ARM. ARM indicator
light on.

4.6 RADAR SIGNAL DETECTING SET AN/APR39(V)2. EH

2. ARM switch - ARM.

The radar signal detecting set indicates the relative position of search radar stations. Differentiation is also made
between various types of search radar and tracking stations.
Audio warning signals are applied to the pilot’s and copilot’s headsets. The radar signal detecting set is fed through
the 50-ampere LH MAIN AVIONICS and RH MAIN
AVIONICS circuit breakers on the copilot’s circuit breaker
panel and protected by the 7.5-ampere APR-39 circuit
breaker on the copilot’s circuit breaker panel. The associated antennas are shown in Figure 3-1. Refer to TM 115841-288-12.

3. FLARE switch (instrument panel) - Press for
each release.

4.6.1 Controls and Function. The operating controls
of the AN/APR-39(V)2 panel (Figure 4-2) are as follows:

4. Dispense button press or mode switch PGRM,
as required.
4.3.11 Flare Operation.

1. FLARE counter - Set for number of flare cartridges in payload module.

NOTE
If the flare detector does not detect burning
of the first flare fired, another flare is automatically fired within 75 milliseconds; if
burning is still not detected, a third and final
flare is fired. If all three flares do not fire,
automatic ejection of flares will stop until
4-4

c. Distance Readout - displays, in kilometers, the distance to a signal emitter selected by the DF operator.

Change 10

CONTROL

FUNCTION

PWR ON

Supplies 28 VDC to the Radar Detecting Set. Fully operational after
one minute.

PWR OFF

Turns system off.

TM 1-1520-237-10

CONTROL
Night-Day filter

AUDIO

HI ALT

PWR
OFF
L

TEST

AA1635_1

FUNCTION
Varies the intensity of the red polarizing face plate filter for day or
night operation. If the IP-1150A is
used in the aircraft, the day-night
switch is not used as the IP-1150A
is night vision compatible.

4.6.2 Processor Unit. The processor has a theater selection switch on the front of the processor. Selection of
one of six theaters is possible depending on mission and
geographical location.

Figure 4-2. AN/APR-39(V)2 Control Panel

CONTROL

FUNCTION

HI ALT

Selects high altitude mode of operation. Selection is based on aircraft mission profile.

LOW

Selects low altitude mode of operation. Selection is based on aircraft
mission profile.

TEST

Initiates system self-test function
when
momentarily
depressed
downward. Permits flight line testing of Radar Detecting Set when
held in the upward position. (Self
test does not test antennas, antenna/
receiver cabling.)

Audio

Controls level of audio output to the
aircraft interphone control system.
Turn to the right for audio volume
increase. Turn to the left for audio
volume decrease.

Direction/Display
(Scope/IP-1150)

Shows alphanumeric symbology
for signals programmed in the processor emitter identification table.

MA indicator

Lamp flashes when low band signals associated with missile guidance systems are correlated with
high band signals associated with
tracking systems.

BRIL control

Varies the brilliance of the alphanumeric symbology.

4.6.3 Self-Test Procedures.
a. The self-test confidence checks all AN/APR-39(V)2
circuits except antennas, high pass filters and detectors in
the high band receivers, bandpass filter and detector in the
low band receiver, high low blanking circuits and antenna/
receiver cabling.
b. The radar detecting set performs a self-test sequence
when the TEST switch on the control unit is set to TEST
and then released. This self-test takes seven seconds, during
which time four different patterns are displayed. The alphanumeric symbology that is displayed at the 12, 3, 6 and 9
o’clock positions will vary depending upon the selected
theater switch on the processor.
c. In the self-test, four patterns will be displayed in sequence (Figure 4-3) on the display MA. Pattern number
one alphanumerics displayed will depend on the selected
theater position on the processor.
d. Observe that the radar signal indicator unit displays
patterns 1-3 and finally, either the NO signal pattern or an
actual threat pattern (Figure 4-3). As each of the first three
patterns are displayed, a different type audio tone will flash
on and off during the display of pattern number 3.
e. An 9H9 symbol will appear in the center of the NO
signal pattern if the control unit HI ALT/LOW switch is in
the HI ALT mode. Conversely, an 9L9 symbol will appear
in place of the 9H9 symbol if this switch is in the LOW
mode.
f. It is important for operators to note if the software
version number displayed at the 12 o’clock position on the
display is the same as the software version sticker attached
to the rear of the processor.

Change 10

4-5

TM 1-1520-237-10

BRI
L

G
NI

PATTERN NO. 1

BRI
L

9
C

G
NI

0
MA

PATTERN NO. 2

PATTERN NO. 3

BRI
L

G
NI

NO SIGNAL PATTERN

AA1635_2
SA

Figure 4-3. Self-Test Patterns AN/APR-39(V)2

4-6

Change 10

TM 1-1520-237-10

4.7 RADAR SIGNAL DETECTING SET AN/APR39A(V)1.
Refer to TM 11-5841-294-12.
4.7.1 Controls and Functions. The operating controls
(Figure 4-4) of the AN/APR-39A(V)1 are as follows:

TEST

PWR
ON
OFF

CONTROL

AUDIO

1
2

FUNCTION

PWR

Controls 28VDC from the No. 1 dc
primary bus.

Locks the switch in the ON position. System is fully operational after approximately one minute. On
power up the synthetic voice will
announce 9APR-39 POWER UP9.
The plus (+) symbol will appear
and be centered on the IP 1150A
cathode ray tube (CRT) during system operation.

OFF

Turns system off. Switch must be
pulled to unlock and turn system
off.

TEST

When momentarily depressed initiates self-test confidence check
(except for antennas and antenna
receiver cabling).

MODE

Selects synthetic voice message
format only. MODE ONE (UP) selects normal voice message format.
MODE TWO (DOWN) selects
test/abbreviated voice message format.

AUDIO

Controls volume to the interphone
system.

Direction/Display
(Scope IP 1150A)

Shows alphanumeric symbology on
a bearing for each processed emitter signal. Does not indicate any
range data.

MA indicator

Not used.

MA switch

Not used.

BRIL control

MODE
+

Varies brilliance of CRT.

AB2422
SA

Figure 4-4. AN/APR-39A(V)1 Control Panel

CAUTION

To prevent damage to the receiver detector crystals, assure that the AN/APR39A(V)-1 antennas are at least 60 yards
from active ground radar antennas or 6
yards from active airborne radar antennas. Allow an extra margin for new, unusual, or high power emitters.
Excessive indicator display brightness
may damage CRT.
4.7.2 Modes of Operation.
a. Self test mode.
(1) After power up, the AN/APR-39A(V)1
synthetic voice will announce 9APR-39
POWER UP9 and the (+) symbol will stabilize in the center of the CRT (Figure 4-5).
Self test should be initiated after approximately one minute. Self test can be performed in MODE ONE or MODE TWO.
In MODE ONE the synthetic voice will
announce 9SELF TEST SET VOLUME, 1,
2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 9. In MODE
TWO the synthetic voice will announce
9SELF TEST SET VOLUME, 5, 4, 3, 2,
19.

Change 10

4-7

TM 1-1520-237-10

(2) The CRT (Figure 4-6) will display specific
software version numbers i.e., operational
flight program (OFP) at the 12 o’clock position and the emitter identification data
(EID) at the 6 o’clock position.
(3) After the software version numbers have
been displayed the test sequence checks the
receivers. A good visual self test will show
two triangles, one at the 6 o’clock and one
at 12 o’clock position on the CRT (Figure
4-6.1). Snowflake symbols (*) will appear
at the 2, 4, 8, and 10 o’clock positions and
will flash if the AN/AVR-2 laser detecting
set is not installed. This is a normal indication and does not effect system performance.
(4) A good self test (no faults detected) ends
with the message 9APR-39 OPERATIONAL9. A bad self test (faults detected)
ends with the 9APR-39 FAILURE9.
b. MODE ONE operation. Selecting MODE
ONE the operator will hear all the normal synthetic voice audio when an emitter has been
processed e. g., the AN/APR-39A(V)1 will announce; 9SA, S-18 12 O’CLOCK TRACKING9. Selection of this mode does not have any
effect on emitters received, processed or displayed, it only affects synthetic voice audio.
c. MODE TWO operation. Selecting MODE
TWO the operator will hear an abbreviated
synthetic voice audio e. g., the AN/APR39A(V)1 will announce; 9MISSILE 12
O’CLOCK TRACKING9.
4.7.3 Function.
a. The radar signal detecting set (RSDS) receives,
processes and displays pulse type signals operating in the C-D and H-M radio frequency
bands. The emitters that it processes and displays are derived from the EID contained in the
user data module (UDM) that is inserted in the
top of the digital processor. In normal circumstances the processor is classified confidential if
a classified UDM is installed.
b. The UDM contains the electronic warfare threat
data that makes up the specific library for a
specific mission(s) or a geographical location
(it is theaterized). When a match of the elec-

4-8

Change 10

AB2423
SA

Figure 4-5. CRT Power Up Display
tronic warfare data occurs the processor generates the appropriate threat symbology and synthetic audio. It is important therefore that the
correct theaterized EID and UDM are installed
for the mission or geographic location.
c. Symbol generation and position relative to the
center of the CRT shows the threat lethality, it
does not show or represent any lethality of
range, but of condition/mode of the emitter.
Highest priority threats (most lethal) are shown
nearest the center. Each symbol defines a generic threat type, symbols are modified to show
change in the status of the emitter. The symbols
are unclassified, the definitions of what the
symbols mean are classified. The complete set
of symbols and definitions are contained in TM
11-5841-294-30-2. Each theaterized library
EID has a specific classified pilot kneeboard
produced with it. The unit electronic warfare
officer (EWO) should contact PM-ASE if sufficient cards are not available within his unit for
the installed EID.
d. The RSDS on specific aircraft has been interfaced with other aircraft survivability equip-

TM 1-1520-237-10

XXX.X

XXX

AB2425

AB2424

Figure 4-6. CRT Version Number Display
ment. The equipment includes the AN/AVR-2
laser detection set, AN/APR-44(V) continuous
wave receiver and the AN/AAR-47 missile
warning system.
4.8 INFRARED COUNTERMEASURE SET AN/ALQ144A(V)1.

WARNING
Do not continuously look at the infrared
countermeasure transmitter (Figure 4-1)
during operation, or for a period of over 1
minute from a distance of less than 3 feet.
Skin exposure to countermeasure radiation for longer than 10 seconds at a distance less than 4 inches shall be avoided.
Ensure the countermeasure set is cooled
off before touching the unit.

Figure 4-6.1. CRT Self Test Display

CAUTION

Observe that the IRCM INOP caution
light illuminates when the OCU ON/OFF
switch is set to OFF. After 60 seconds, observe that the IRCM INOP caution light
extinguishes.
The countermeasure system provides infrared countermeasure capability. The system transmits radiation modulated mechanically at high and low frequencies using an
electrically-heated source. A built-in test feature monitors
system operation and alerts the pilot should a malfunction
occur. The system is made up of a control panel on the
instrument panel and a transmitter on top of the main rotor
pylon aft of the main rotor. On helicopters Serial No. 7822987 and subsequent, the countermeasure system functionally interfaces with the caution/advisory warning system through the left relay panel. The countermeasure
system gets dc electrical power from the No. 2 dc primary

Change 10

4-9

TM 1-1520-237-10

circuit breaker panel and the No. 2 junction box. The 28
vdc is routed through the IRCM PWR circuit breaker in
the No. 2 junction box to the transmitter. The No. 2 dc
primary bus also supplies 28 vdc through the IRCM
CONTR circuit breaker on the No. 2 dc primary circuit
breaker panel to the control unit. Panel lighting of the control unit is controlled by the INSTR LTS NON FLT control on the upper console. When the control unit ON-OFF
switch is placed ON, the power distribution and control
circuits are activated and the ON lamp is lit for about 60
seconds on helicopters prior to Serial No. 78-22987. The
source begins to heat, the servo motor and drive circuits are
energized, turning on the high and low speed modulators,
and a signal is applied to stabilize system operations before
energizing the built-in test function. After a warmup period
the stabilizing signal is removed, and the system operates
normally. Placing the ON-OFF control switch momentarily
to OFF causes the power distribution and control circuits to
de-energize the source and initiates a cooldown period.
During the cooldown period, the servo motor drive circuits
remain in operation, applying power to the motors that
cause the modulators to continue turning. The INOP light
will remain on during cooldown cycle, and on helicopters
Serial No. 78-22987 and subsequent, the IRCM INOP caution light will be lit. After the cooldown period, the power
distribution and control circuits de-energize, all system operating voltage is removed and the IRCM INOP caution
light, or the INOP light, will go off. On helicopters prior to
Serial No. 78-22987, if a malfunction occurs during system
operation, the INOP light on the control unit will go on and
the cooldown period will automatically begin. On helicopters Serial No. 78-22987 and subsequent, if a system malfunction causes the IRCM INOP caution light on the caution panel to go on, the IRCM INOP caution light will
remain lit until the control panel ON-OFF switch is momentarily placed OFF. The system can be returned to operating mode by momentarily placing ON-OFF switch
OFF, then ON, provided the cause of the malfunction has
cleared. For additional information, refer to TM 11-5865200-12.

4.8.2 Controls and Function. Controls for the AN/
ALQ-144 are on the front panel of the control unit. The
function of each control is as follows:

4.8.1 Infrared Countermeasure System Control
Panel. Control of the countermeasure set is provided by
the operator control panel on the helicopter instrument
panel. On helicopters prior to Serial No. 78-22987, the control panel has one switch (ON-OFF) and an ON indicator
light. On helicopters Serial No. 78-22987 and subsequent,
only a power ON-OFF switch is on the control panel.
Power to operate the countermeasure set is supplied from
the No. 2 dc primary bus through a circuit breaker, marked
IRCM CONTR.

3. ON-OFF switch - ON (helicopters Serial No.
78-22987 and subsequent).

4-10

Change 10

CONTROL

FUNCTION

ON-OFF switch

Turns set on and off.

ON indicator light
(green) (Helicopters
prior to Serial No.
78-22987)

Indicates system is in a 45 to 75
second warmup mode.

INOP indicator light
(red)
(Helicopters
prior to Serial No.
78-22987)

Indicates malfunction has occurred
or countermeasure system is in
cooldown cycle.

IRCM INOP caution light (Helicopters Serial No. 7822987
and
subsequent)

Indicates malfunction has occurred
or the countermeasure system is in
cooldown cycle.

4.8.3 Operation.
1. ON and INOP PRESS TO TEST indicator
light - Press. Indicator light should go on. (On
helicopters prior to Serial No. 78-22987.)
2. ON-OFF switch - ON. Green indicator light
should light for 45 to 75 seconds, then go off.
(On helicopters prior to Serial No. 78-22987.)
NOTE
If the INOP indicator or IRCM INOP caution light goes on after the ON indicator (helicopters prior to Serial No. 78-22987) goes
off, place the power switch OFF.

4.8.4 Stopping Procedure.
ON-OFF switch - OFF. The transmitter will
continue to operate for about 60 seconds during
the cooldown cycle. INOP indicator or IRCM
INOP caution light as applicable should remain
on during cooldown cycle.

TM 1-1520-237-10

4.9 ECM ANTENNA SWITCH.

The ECM antenna switch is a three-position switch on
the instrument panel (Figure 2-9), providing control of
ECM antenna deployment and retraction. The switch is
spring-loaded to center (OFF), with positions marked EXTEND and RETRACT. Normal operation of the switch is
as follows:

CAUTION

The ECM antenna can be extended with
the helicopter on the ground if the radar
altimeter is turned off or removed from
the installation, or the L (LO SET) indicator is set below the radar altimeter indication.
a. Extend. When the helicopter is on the ground with all
systems working properly and the radar altimeter L (LO
SET) indicator is set above the radar altimeter indication,
the antenna cannot be extended because of the interlock
system. When the helicopter is in flight, with the copilot’s
radar altimeter indication above the L (LO SET) indicator,
the antenna can be extended until it reaches the fully extended position, by momentarily placing the switch to EXTEND. Once the extension or retraction process has started,
it cannot be overridden with another command from the
ECM ANTENNA switch. The cycle can be interrupted by
turning off the radar altimeter or setting L SET bug above
radar altitude indication. When the antenna is fully extended, a light on the ECM operator’s console marked AN-

TENNA DEPLOYED, will go on. The ANTENNA EXTENDED caution light on the caution/advisory panel will
not go on when the antenna is extended. It is a condition
light rather than an antenna position light.
NOTE
Automatic ECM antenna retraction is controlled by the copilot’s radar altimeter L (LO
SET) indicator when the altimeter is turned
on.
b. Retract. If the antenna is extended, the pilot may momentarily select RETRACT to return the antenna to the
retracted position. The antenna will automatically retract if
the helicopter descends below the altimeter L indicator setting or a failure occurs in the radar altimeter. When the
antenna is fully retracted, the ANTENNA RETRACTED
status advisory light on the caution/advisory panel will go
on and remain on as long as the antenna stays in that position. The ANTENNA DEPLOYED and ANTENNA EXTENDED lights should be off with antenna retracted. An
emergency retract switch accessible to the ECM operator
may be used to retract the ECM antenna if a failure occurs
in the cockpit retract system. A light next to the switch
indicates when the antenna is extended.
c. Emergency ECM Antenna Retract Switch. An emergency ECM antenna retract switch on the antenna relay
assembly on the ECM equipment rack, provides a backup
mode of retraction of the antenna if a failure occurs in the
cockpit ECM ANTENNA switch. To retract, the switch
must be held at up until the antenna is fully retracted and
the ANTENNA RETRACTED advisory light is on.

Pages 4-12 through 4-12.2 deleted.
Change 10

4-11/(4-12 Blank)

TM 1-1520-237-10

4.10 COUNTERMEASURES SET AN/ALQ-156(V)2.
EH

Countermeasure set AN/ALQ-156(V)2 consists of Receiver Transmitter RT-1220. Control indicator C-10031,
and four each circular horn antenna AS-3650. Antenna locations are illustrated on Figure 3-1. The countermeasures
set provides aircraft protection against infrared-seeking
missiles by detecting valid targets and sending pulses to
flare dispenser M130. Decoy flares are then launched away
from the aircraft. Power to operate the countermeasures set
is taken from the No. 1 dc primary and No. 1 ac primary
buses through circuit breakers located on the copilot’s circuit breaker panel (Figure 2-20).
4.10.1 Basic Principles of Operation. Incoming and
outgoing RF signals are routed between the circular horn
antennas and the receiver transmitter through coaxial
cables. The transmit signal is modulated and amplified in
the receiver transmitter, and routed alternately to forward
and aft antennas. After each pulse transmission, return signals are received by the same antenna used for transmission
and routed to the receiver transmitter for processing. When
an approaching missile is detected, the countermeasures set
sends a pulse to flare dispenser M130. If armed, the flare
dispenser launches a decoy flare to draw the missile away
from the aircraft.
4.10.2 Controls, Displays, and Functions.

WARNING
During take-off, landing and ground operations, the ALQ-156 POWER switch
must be in the OFF position. Failure to
comply may cause inadvertent release of
flares resulting in personal injury or damage to equipment.
Control indicator C-10131, located on the instrument
panel (Figure 2-9) to the right of the ECM ANTENNA
switch, contains controls and status indicators for system
operation. The control indicator front panel is illuminated
by integral lighting. Controls and indicator of C-10131 are
shown in Figure 4-1 described below:

CONTROL/
INDICATOR

FUNCTION

POWER switch

Places countermeasure set in
operate (ON) mode. The switch is a
positive-locking type and cannot be
accidentally shut off. The switch
must be pulled out and then down
to turn off the countermeasure set.

TEST FLARE
switch

Tests ALQ-156/M130 systems and
enables test flare launch in ON
condition. The switch is a
momentary depress-release type. It
should be used only in conjunction
with flare dispenser M130 test
procedures.

PUSH FOR
STANDBY
STATUS
pushbutton

Places countermeasure set in
standby or operate mode. When
depressed, switch places system in
standby mode and upper half of
indicator shows STBY. In 9out9
position, switch places system in
operate mode. During initial
warmup, lower half of indicator
shows WRMUP. After warmup, the
indicator is blank to show that the
system is on and ready for flare
dispense.

4.10.3 Operation. The following procedures shall be
followed to operate the countermeasure set:

WARNING
Do not stand within six feet of Aircraft
Survivability Equipment (ASE), ALQ-156
and ALQ-162, transmit antennas when
the ASE equipment is on. High frequency
electromagnetic radiation can cause internal burns without causing any sensation
of heat.
When ALQ-156 POWER switch is ON,
and the M130 Flare system is armed, a
flare can be fired.
1. M130 flare system ARM/SAFE switch SAFE.

4-13

TM 1-1520-237-10

NOTE
Prior to beginning the turn-on procedure, ensure that the push for standby pushbutton is
in the 9out9 position (not depressed).
2. ALQ-156 POWER switch - ON. Observe that
status indicator shows WRMUP, indicating that
receiver transmitter is in warmup mode.
NOTE
The actual length of time that the WRMUP
lamp remains on depends upon a combination of equipment operating status and environmental temperature. Under normal operating conditions, and with air temperature
about 77°F (25°C), WRMUP lamp will go
out in approximately 8 to 10 minutes.
3. Observe that WRMUP lamp goes out, indicating that the receiver transmitter is now in the on
condition.
4. PUSH FOR STANDBY/STATUS pushbutton
- Push once to place the countermeasure set in
standby. Subsequent depressions switch the
countermeasure set alternately from on to
standby.

4.11.1 Basic Principles of Operation. Incoming signals received from SAM and AIM missiles using Continuous Wave (CW) for guidance are validated by the countermeasures set. Depending upon validation results, the system
initiates jamming action and/or warns the crew of approaching missiles. Automatic jamming/warning decisions are determined by warning and jamming thresholds preprogrammed in the system. The countermeasures set may
be used in stand-alone fashion or in conjunction with AN/
APR-39(V)2 Radar Warning Receiver (RWR). The RWR
processes and displays threat information. A Built-In Test
(BIT) automatically and continually tests systems operations. Malfunctions cause a no-go lamp to light in the control unit front panel. The countermeasures set can be structured to counter different threats by programming the
program module assembly in the front of the receiver transmitter. The programming is done before flight by the
ground crew, as the receiver transmitter is not within operator reach.
4.11.2 Controls, Displays, and Functions. Located
in the center of the lower console, Control Unit C-11080
contains controls and indicators necessary for countermeasures set operation. The control unit is described below:

CONTROL/
INDICATOR
VOLUME control

Controls tone generator volume. A
tone is generated in the aircraft
headset immediately upon threat
detection.

BIT test switch

Initiates automatic and continuous
Built-In-Test of countermeasures
set operations.

Lamp test switch

Tests lamp functions of WRMUP,
NO GO lamps.

Function switch

Controls
countermeasures
set
operation. OFF removes power
from the set. STDY provides
warmup power but does not enable
transmit-receive circuits. RCV
turns on the receiver for
maintenance testing of antenna,
sensing, and processing circuits.
OPR provides full operational
power to both receiver and
transmitter.

5. M130 flare system ARM/SAFE switch - ARM
(Figure 4-1).
4.10.4 Stopping Procedure. The following procedure
shall be used to turn off the countermeasures set:

FUNCTION

ALQ-156 POWER switch - OFF.
4.11 COUNTERMEASURES SET AN/ALQ-162(V)2.
EH

Countermeasures set AN/ALQ-162(V)2 consists of Receiver Transmitter RT-1377, Control Unit C-11080, and
two each antenna AS-3554. Antenna locations are illustrated on Figure 3-1. The countermeasures set provides
warning and protection against surface-to-air (SAM) and
airborne interceptor missiles (AIM). Missile radar signals
are detected by the system, modulated internally, and retransmitted as false, misleading echoes. Power to operate
the countermeasures set is taken from the DC MON and
No. 2 ac primary buses through circuit breakers located on
the pilot’s circuit breaker panel (Figure 2-20), refer to TM
11-5865-229-10.

4-14

TM 1-1520-237-10

CONTROL/
INDICATOR
Warm up & No Go
Lamps

FUNCTION
Indicates countermeasures set status. WRMUP appears when unit is
first turned on and goes out after
approximately 3 minutes. NO GO
will light if BIT operation detects a
system failure.

warmup period is required. Do not attempt
operation of the unit until warmup is successfully completed.
3. WRMUP lamp - Check that lamp goes out after 3 minutes.

WARNING

4.11.3 Operation.

WARNING
When the countermeasures set is operating, electromagnetic radiation is present.
DO NOT OPERATE if personnel are
within six feet of transmit antennas. High
frequency electromagnetic radiation can
cause internal burns without causing any
sensation of heat.
NOTE
A complete operational test consisting of a
lamp test, operator-initiated BIT test, and a
signal test is incorporated in normal operational procedures. The test shall be performed before flying any mission that requires use of the countermeasures set.
The following procedures shall be used to operate the
countermeasure set under usual conditions: OPERATORINITIATED BIT TEST
NOTE
Before beginning step 1, turn control unit
VOLUME control fully clockwise.
1. Control unit function switch - STBY. Observe
front panel and WRMUP lamps light. A tone
should be heard briefly in the headset.
2. Lamp test switch - Press and observe all four
lamps light in pushbutton switch.
NOTE
If the countermeasures set has been without
power for 30 seconds or more, a 3 minute

The countermeasure set will radiate powerful, high-frequency electromagnetic energy when countermeasures set function
switch is set to OPR. Ensure personnel are
at least six feet from antennas while countermeasures set is in operate mode.
4. Control unit function switch - OPR.
5. BIT switch - Depress. A tone should be heard
in the headset.
4.11.4 Stopping Procedure. The following procedure
shall be used to turn off the countermeasures set: Control
unit function switch - OFF.
4.12 HEADS UP DISPLAY AN/AVS-7.
Heads up display (HUD) AN/AVS-7; (Figure 4-7) consists of signal data converter CV-4229/AVS-7 (SDC) located in the avionics compartment, the converter control
C-12293/AVS-7 (CCU) located on the lower console, and
the display, SU-180/AVS-7 (DU) consisting of the optical
unit (OU) and power supply calibration unit (PSCU). Two
thermocouple amplifiers are located in the avionics compartment and two HUD control switches are located on the
pilot’s collective sticks. The HUD system serves as an aid
to pilots using the AN/AVS-6 (ANVIS) during night flight
operations by providing operational symbology information
about the aircraft. There are two programming modes and
one operational mode which allow both pilots to independently select the symbology for their respective display
modes from a master set of symbols in the signal data converter. Power to operate the HUD system is provided by the
26V ac essential bus and the 28V dc bus through circuit
breakers marked HUD REF and HUD SYS.
4.12.1 Basic Principles of Operation. The pilots can
independently select from four normal symbology modes
and four declutter modes that were pre-programmed. Declutter mode has four vital symbols that will always be
displayed: Airspeed, Altitude (MSL), Attitude (pitch and
roll), and Engine Torque(s). An adjust mode, during opera-

Change 4

4-15

TM 1-1520-237-10

FOCUS
RING

E T
Y C
E E
L
E
S

OPTICAL
UNIT

POWER SUPPLY
AND CALIBRATION
UNIT
L / R EYE
SELECT

DISPLAY UNIT
SU−180/AVS−7

CONVERTER
CONTROL
C−12293 / AVS−7

BRT

D/U

DIM

L/R

FAIL

DSPL POS

BIT

ALT / P / R

ACK

SEL
PGM

ADJ

1−4

INC

DEC
NXT
DCLT

DIM

MODE

OP
CP−PGM

L/R

PLT

+
MODE
1−4

BRT

DSPL POS

CPLT

P−PGM

D/U

DCLT

OFF

AA9221A
SA

Figure 4-7. Heads Up Display AN/AVS-7

4-16

TM 1-1520-237-10

tion, is used to adjust barometric altitude, pitch, and roll. If
the HUD system loses operating power after adjustments
have been made, the brightness, mode, barometric altitude,
pitch, and roll must be adjusted as necessary. The system
self test is divided into power-up or operator initialized
built-in-test (BIT) and in-flight BIT. The system BIT is
initialized during power-up or selected by the operator. Part
of the BIT is a periodic test that is performed automatically
along with normal system operation. A failure of the SDC,
or the pilot’s DU will illuminate the CCU FAIL light and
display a FAIL message on the display unit. When a FAIL
message is displayed on the DU, the operator should acknowledge the failure and re-run BIT to confirm the fault.
4.12.2 Controls and Functions. The CCU, located on
the lower console, (Figures 2-8 and 4-7), and the control
switches on the pilot’s collective stick (Figure 2-14) are
controls and indicators necessary for HUD operation. The
EYE SELECT L/R position is set when display units are
connected prior to operation. A focus ring on the OU provides control for focusing the display. The OU is adjusted
by the manufacturer and under normal conditions adjustment is not required.

CONTROL/
INDICATOR

FUNCTION

ADJ/ON/OFF

Selects adjust mode, enabling the
INC/DEC switch to adjust altitude,
pitch, or roll. Turns power on or off
to HUD system.

P-PGM/OP/CPPGM

Selects pilot program mode,
operational mode, or copilot
program mode. Used with the
PGM NXT/SEL switch.

BIT/ACK

Selects built-in-test or used to
acknowledge a displayed fault,
completion of an adjustment, or
completion of a programming
sequence.

ALT/P/R
INC

DEC/

a. The converter control is described below:

CONTROL/
INDICATOR

FUNCTION
PGM NXT/SEL

CPLT
BRT/DIM

Copilot’s control
brightness.

for

display

DSPL POS D/U/
L/R

Copilot’s control for display
position down/up (outer knob) and
left/right (inner knob).

MODE 1-4/
DCLT

Copilot’s mode select 1-4 and
declutter switch.

Active when adjust mode is
selected
to
decrease/increase
altitude/pitch roll. When adjusting
altitude (MSL) a momentary
movement of the INC/DEC switch
will change data in 5 feet
increments. When the INC/DEC
switch is held for one second data
will change 10 foot increments.
Pitch and roll change in increments
of one degree.
Active when program mode is
selected. Allows operator to preprogram the four normal modes
and four declutter modes. Operator
can select a flashing symbol for
display and/or go to the next
symbol. Once complete, operator
toggles the ACK switch to save
programmed display.

b. Pilot’s collective controls are described as follows:

PLT
BRT/DIM

Pilot’s
control
brightness.

for

display

DSPL POS D/U/
L/R

Pilot’s control for display position
down/up (outer knob) and left/right
(inner knob).

MODE 1-4/
DCLT

Pilot’s mode select
declutter switch.

1-4

FAIL

Indicates a system failure.

Indicates system ON.

and

CONTROL

FUNCTION

BRT/DIM

Allows pilot’s to control brightness
of their respective displays.

MODE/DCLT

Allows pilot’s to select respective
display modes or declutter modes.

c. Attach optic unit to either ANVIS monocular housing. Set EYE SELECT switch on PSCU to L or R.

4-17

TM 1-1520-237-10

4.12.3 Modes of Operation. There are two programming modes and one operational mode for the HUD system
selected by the programming switch on the CCU. The adjust mode is a submode under the operational mode.

cated on the lower console. Modes are defined by selecting
from a master symbology menu (Figure 4-8 and Table 4-1).
Up to eight display modes, four normal and four declutter,
can by programmed for each user and can be selected for
display using the display mode selection switch on the pilot’s collective control or on the CCU. The default declutter
mode has a minimum symbology display of:

1. Pilot programming switch - Set to P-PGM.
2. Copilot programming switch - Set to CP-PGM.
3. Operation (flight mode) switch - Set to OP.
(Adjust - ADJ/ON/OFF switch to ADJ).
4.12.4 Display Modes. Symbology display modes are
programmable by the pilots via the converter control lo-

•

Airspeed - No. 25.

•

Altitude (MSL) - No. 7.

•

Attitude (pitch and roll) - No. 1, 5, 6, 20, 26.

•

Engine Torque(s) - No. 22, 23.

Table 4-1. UH-60A/L Master Mode Symbology Display (HUD)
No.

Symbol

Source

Range/Description

Angle of Pitch Scale

HUD System

6 30° (10° units, tic marks flash when angle of pitch is >
6 30°).

Bearing to Waypoint Pointer

Doppler

0 - 359° (cursor will invert 9V9 when aircraft is moving
away from waypoint).

Compass Reference Scale

HUD System

0 - 359° (10° units).

Aircraft Heading Fix Index

HUD System

Fixed Reference Mark.

Angle of Roll - Pointer

Copilot’s Vertical
Gyro

6 30° (right turn moves pointer to right, pointer flashes >
6 30°).

Angle of Roll - Scale

HUD System

6 30° (10° units).

Barometric
(MSL)

Air Data System

-1000 to 20,000 feet (set during adjustment mode).

Adjust/Program
Message

HUD System

ADJ or PROG.

OK/FAIL

HUD System

OK or FAIL.

Velocity Vector

Doppler

0 - 15 knots/15 kilometers, 0 - 359°.

Rate of Climb Pointer

Air Data System

6 2000 feet-per-minute (used with vertical speed scale,
No. 15).

Radar Altitude (AGL) Numeric

Pilot’s Radar Altimeter

0 - 1000 feet (0 - 200 feet, 1 foot units; 200 - 1000 feet,
10 foot units; disappears above 999 feet, and reappears
below 950 feet).

Minimum Altitude Warning

Pilot’s Radar Altimeter

Blinking square around symbol - No. 12, (set on pilot’s
low warning bug).

Radar Altitude
Analog Bar

Pilot’s Radar Altimeter

0 - 250 feet (disappears at 250 feet, reappears at 230 feet;
digital readout symbol, No. 12).

AGL, Vertical Speed Scale

HUD System

0 - 200 feet/6 2000 feet-per-minute.

HUD Fail Message

HUD System

CPM, SDR, SDA, PS, PDU, CPDU, NAV, PGM; can
be cleared from the display by selecting ACK (see note).

4-18

Altitude
Mode

(AGL)

TM 1-1520-237-10

Table 4-1. UH-60A/L Master Mode Symbology Display (HUD) (Cont)
No.

Symbol

Trim (Slide Ball)

MST, MEM,
Messages

Source
SAS/FPS
puter
HOOK

Master
Panel

Com-

Range/Description
6 2 balls (left/right).

Caution

MST, MEM, HOOK cannot be cleared from the display
by selecting ACK.

Sensor, Engine, Fire, RPM
Warnings

Master Warning
Panel

ATT, ENG 1 or 2, FIRE, RPM; ATT can be cleared
from the display by selecting ACK (see note). ENG,
FIRE, and RPM cannot be cleared.

Horizon Line (pitch, roll)

Copilot’s Vertical
Gyro

Pitch: 6 30°

Display Mode Number

HUD System

1N - 4N for normal modes, 1D - 4D for declutter modes.

Torque Limits

Torque
ducer

Trans-

0 - 150%
Yellow (>100%), (solid box)
Red (>110%) Thresholds (solid box flashes).

Torque - Numerics

Torque
ducer

Trans-

0 - 150% (flashes when engine torque separation is greater
than 5% threshold) Maximum % torque split between
cockpit panel and HUD is 3%.

Ground Speed

Doppler

Indicated Airspeed

SAS/FPS
puter

Attitude Reference Indicator

HUD System

Represents helicopter.

Engines Temperature

Thermocouple
Amplifers

0 - 999°C (0 - 755°C - 999°C, 1° units) Maximum split
between cockpit and HUD is 6 15°.

Distance to Waypoint

Doppler

0 - 999.9 km.

Bearing to Waypoint - Numeric

Doppler

0 - 359°

Roll: 0 - 359°.

0 - 999 knots/0 - 530 km/h (dependent on doppler).
Com-

30 - 180 knots (no symbol 30 knots and below, reappears
at 32 knots).

NOTE: After ACK is used to acknowledge a
fault, the fault will not reappear until BIT is
selected or power is cycled off and on.
4.12.5 Operation.
4.12.5.1 Starting Procedure.

WARNING

1. ADJ/ON/OFF switch - OFF.
2. Optical unit support clamps - Installed on ANVIS. Verify clamps can by rotated.

Failure to remove the ANVIS neck cord
prior to operation of the HUD may prevent egress from the aircraft in an emergency.

NOTE
4. ANVIS neck cord - Removed.
Check surface of lens for cleanliness. Clean
in accordance with TM 11-5855-300-10.
3. DU lens - Check.

5. Optical unit - Install on ANVIS. Attach optical
unit to either monocular housing. Do not
tighten OU clamp completely with thumbscrew

4-19

TM 1-1520-237-10

at this time. The OU (display) may have to be
rotated to horizon after the system is operating.
NOTE
The helmet may now have to be rebalanced.
6. EYE SELECT switch on PSCU - L or R.

WARNING
CCU ADJ/ON/OFF switch must by OFF
before connecting or disconnecting quickrelease connector.

CAUTION

The AN/AVS-7 system should not be used
if the quick-release connector is not in
working order.
7. PSCU - Connect. Connect PSCU to quickrelease connector by rotating the connector engagement ring.

CAUTION

Keep the protective caps on the ANVIS
whenever it is not in use. Operate the ANVIS only under darkened conditions.
NOTE
Ensure ANVIS operator procedures have
been completed.
8. P-PGM/OP/CP-PGM switch -OP.
9. ADJ/ON/OFF switch - ON. System ON and
FAIL lights illuminate and BIT will initiate automatically.
10. FAIL light - Check. Light should go out after
ten seconds. BIT is complete.
NOTE
Allow one minute for display warm-up. Display intensity is preset to low each time

4-20

Change 4

ADJ/ON/OFF switch is set from OFF on
ON.
If a fault is displayed in the DU, acknowledge fault and re-run BIT to confirm fault.
11. BRT/DIM switch - As desired.
12. DSPL POS control - As required. Center display in field of view.
13. Display aligned to horizon - Check. Tighten
OU clamp.
4.12.5.2 Operator Self Test (BIT).
1. BIT/ACK switch - Press to BIT and hold. The
ON and FAIL light will illuminate. At end of
BIT, FAIL indicator will extinguish.
2. BIT/ACK switch - Release.
4.12.5.3 Displayed System Faults. The system self
test is divided into power-up or operator initialized builtin-test (BIT) and inflight BIT. The faults result as warnings
and messages that blink at a rate of two per second in the
display units. Part of the BIT is a periodic test that is performed automatically along with normal system operation.
This BIT monitors and/or tests SDC functions and/or signals. A failure of the SDC, NAV signals pilot’s DU, will
illuminate the converter control FAIL light and display a
FAIL message CPM, SDR, SDA, PS, NAV, PDU or
CPDU on the display unit. An attitude (ATT) sensor indication will be displayed when a gyro invalid condition exits. ATT, NAV, PDU, CPDU, and all SDC faults can be
cleared by setting BIT/ACK switch to ACK. The following helicopter status messages are also displayed.
1. The caption MST (first priority) indicates operation of the master caution warning lamp.
This message will disappear during the rest of
the main warning lamp operation.
2. The caption MEM (second priority indicates
that the doppler data is not updated. a previous
computed data is available. This message will
appear simultaneously with the MEM lamp on
the doppler operating panel.
3. The caption HOOK (third priority) indicates
the cargo hook is open.

TM 1-1520-237-10

3
2

5
6

26
25

150
12.3
736T1
710T2

1630B

PROG
OK

9
10

102A
71G

100

11
92

12
ATT ENG1 FIRE RPM

MST

CPM

22
21

15
20
19

AA9222
SA

Figure 4-8. Master Mode Display
The message will appear simultaneously with
the indication lamp in the cockpit.
Setting BIT/ACK switch to ACK will not clear
MST, MEM, or HOOK status messages from the DU.
Engine, FIRE and RPM warnings cannot be cleared from
the DU. The faulty unit or warning must be removed from
the aircraft. When both engines fail at the same time, engine priority is: ENG 1 then ENG 2.
4.12.5.4 Programming Procedure.
NOTE
The programming procedure for the pilot
and copilot is identical except for the location of controls on the CCU.
1. Select mode to be programmed (1N-4N). The
first mode that will appear is 1N (normal mode
1).
2. P-PGM/CP-PGM/OP switch - P-PGM or CPPGM.
3. PROG blinking in display - Check. Verify that
a complete set of symbology is displayed and

attitude reference symbol is blinking. Verify
PGM is displayed in the HUD FAIL message
location for the DU not being programmed.
4. BIT/ACK switch - ACK to program the full
display or go to step 5 and select desired symbols.
5. PGM SEL/NXT control - SEL to select symbol. Selected symbol stops blinking. If symbol
is not desired, toggle switch to NXT and the
symbol will disappear.
NOTE
All symbols have been programmed when
the PROG annunciator is the only symbol
flashing.
6. BIT/ACK switch - ACK. (Hold switch to
ACK for one second.)
7. OK displayed - Check. (OK will be displayed
for two seconds.)

4-21

TM 1-1520-237-10

A
K

B
L

C
M

D
N

E
O

F
P

G
Q

H
R

1 2

3 4 5 6

I
S

J
T

7 8 9 0

%
TEST

S.G.
VER X.XX

DD / MM / YYYY

UH−60

NOTE
VERSION NUMBER AND DATE WILL
CHANGE AS SOFTWARE IS UPDATED.

AA9223
SA

Figure 4-9. Symbol Generator Test Mode
NOTE
If programming is not accepted, FAIL will
be displayed. If a FAIL message is displayed, attempt to reprogram the same
mode, if FAIL reappears notify maintenance.
Declutter mode is recognized by flashing
ground speed indicator in lieu of attitude reference symbology.
8. MODE 1-4/DCLT - DCLT (1D-4D). The first
DCLT mode that will appear is 1D (declutter
mode 1).
NOTE
If MODE 1-4/DCLT switch is toggled to
DCLT a second time the display will cycle
back to the DCLT’s normal mode (1N-4N).
The MODE 1-4/DCLT switch must be set
to MODE 1-4 to advance to another normal
mode.
9. Repeat steps 4 through 7, for declutter.

4-22

10. MODE 1-4/DCLT switch - As required.
11. Repeat steps 4 through 10 until all desired
modes are programmed.
12. P-PGM/CP-PGM/OP switch - OP.
4.12.5.5 Adjustment of Barometric Altitude, Pitch,
and Roll.

WARNING
An improperly adjusted barometric altimeter will result in an improperly set
HUD barometric altitude display.
NOTE
Barometric altimeter should be set to the
most current altimeter settings, field elevation.
1. Ensure P-PGM/CP-PGM/OP switch is in the
OP position.

TM 1-1520-237-10

2. ADJ/ON/OFF switch - Pull and set to ADJ.
3. ADJ blinking in display - Check.
NOTE
Changes to barometric altimeter settings require a corresponding change to the HUD
barometric altitude. Each .01 inch change in
pressure equals 10 feet.
4. INC/DEC switch - As required.

4.12.5.7 System Shutdown Procedure.
1. ADJ/ON/OFF switch - OFF.
2. Turn off ANVIS.

WARNING
CCU ADJ/ON/OFF switch must be OFF
before connecting or disconnecting quickrelease connector.

5. BIT/ACK switch - ACK.
6. Repeat steps 3 through 5 for pitch and roll.
7. ADJ/ON/OFF switch - ON.

Do not disconnect DU by pulling on the
cable connected to the PSCU. The DU
could be damaged or the cable may separate from the PSCU creating an explosive
atmosphere hazard.

4.12.5.6 In-flight Operation.

WARNING
Whenever the symbology displayed in the
DU is suspected of being incorrect the pilot’s will compare the data with the aircraft instrument indicator and take the
appropriate action.
Excessive brightness of the symbology display may impair vision outside the cockpit.
Interruption of electrical power, such as
change over from APU generator to NO.
1 and NO. 2 generators and vice versa,
will cause DU to default to dim and
MODE 1N. Any adjustments made to the
barometric altitude, pitch and roll prior
to flight will be lost, thereby decreasing
the accuracy of the barometric altitude,
pitch and roll.
1. BRT/DIM switch - As desired.
NOTE
Whenever the symbology is interfering with
the outside visibility, decluttering may be
selected to remove symbology.
2. MODE 1-4/DCLT switch - As desired.

Do not attempt to egress the aircraft without performing disconnect as this may result in neck injury.

CAUTION

Do not disconnect DU by pulling on the
cable. To do so may damage the DU.
3. Display unit - Disconnect. Disconnect DU by
grasping the PSCU and rotating the quickrelease connector engagement ring and pull
downward. Remove OU and remove from the
ANVIS and place into storage case.
4. Reattach neck cord to ANVIS.
4.12.5.8 Emergency Egress. The quick-release feature allows you to exit quickly from the aircraft in an emergency without:
a. Damaging or turning the unit off.
b. Getting tangled in cords.
c. Being restrained in the cockpit by hardwired
connections.
d. Removing ANVIS.
It is up to the operator to determine the desired mode
of disconnect based upon his evaluation of the emergency

Change 1

4-23

TM 1-1520-237-10

condition and whether or not the ANVIS goggles will be
needed following egress. The available means of disconnect are as follows:

CONTROL/
FUNCTION

INDICATOR

AN-ALQ-156
a. Release the ANVIS goggles from the helmet.
CM JAM
b. Disconnect the OU from the ANVIS goggles
via the thumbscrew.

CM INOP

c. Grasp PSCU and pull down.
4.13 ASE STATUS PANEL.

Lights if the AN/ALQ-156 is being
jammed.
Lights if the AN/ALQ-156 R/T
fails self-test.
NOTE

The ASE status panel (Figure 4-10) is designed to integrate several ASE indicator lights (for various ASE systems
installed) into one location and also tie those status lights to
the MASTER CAUTION light and the caution/advisory
panel. The status panel provides status lights for three ASE
systems currently designated to be installed, and two blank
slots for expansion.

Only a NO GO light will trip the
ASE caution light on the caution/
advisory panel and the MASTER
CAUTION light.
AN/ALQ-144
IRCM INOP

Lights if the AN/ALQ-144 R/T
fails self-test.
NOTE

CONTROL/
FUNCTION

INDICATOR

NO GO

Lights if the AN/ALQ-162 R/T
fails self-test.

CW THRT

Lights if the AN/ALQ-162 detects
a continuous wave (CW) threat
radar.

CW JAM

Lights if the AN/ALQ-162 is being
jammed.
NOTE

Only a NO GO light will trip the
ASE caution light on the caution/
advisory panel and the MASTER
CAUTION light.

4-24

This condition will trip both ASE
caution light on the caution/
adivisory panel and the MASTER
CAUTION light.

TM 1-1520-237-10

ALQ
162

CW
THRT

NO
GO

CW
JAM

ALQ−156

CM
JAM

CM
INOP

ALQ−144

IRCM
INOP

AA1305
SA

Figure 4-10. ASE Status Panel

4-25

TM 1-1520-237-10

Section II ARMAMENT
4.14 ARMAMENT SUBSYSTEM.
The subsystem is pintle-mounted in each gunner’s window at the forward end of the cabin section (Figure 4-11).
The two M60D 7.62 millimeter machineguns are freepointing but limited in traverse, elevation, and depression
field of fire. Spent cartridges are collected by an ejector
control bag on the right side of the weapon. An ammunition
can assembly is on the left side, refer to TM 9-1005-22410. For information on the gun mount, refer to TM 9-1005262-13.
4.15 MACHINEGUN 7.62 MILLIMETER M60D.
The machinegun (Figure 4-12), is air-cooled, gasoperated and automatic. It uses standard 7.62 mm ammunition (Table 4-2). Headspace is fixed to permit quick
change of barrels. Designed primarily for operation in the
air, the M60D has an aircraft ring-and-post sighting system.
The weapon is pintle-mounted and is held by a quickrelease pin. The weapon mount is on a rotating arm assembly which allows the weapon to be locked outboard in the
firing position, or stowed inside the aircraft when the rotating arms are locked in the inboard position. It is easily
removed from the helicopter and can be used for ground
defense with the bipod extended. For more detail of the
M60D, refer to TM 9-1005-224-10.
4.15.1 Controls. Controls for the M60D are on the
weapon and consist of: barrel lock lever, safety, cocking
handle, cover, latch grip and trigger and magazine release
latch.
4.15.1.1 Barrel Lock Lever. The lever (Figure 4-12) is
at the right front of the receiver. It is attached to the barrel
locking shaft and turns to lock or unlock the barrel assembly.
4.15.1.2 Safety. The safety (Figures 4-12 and 4-13), at
the lower front of the receiver, consists of a cylindrical pin
with a sear clearance cut which slides across the receiver to
block the sear and prevent firing. The ends of the pin are
marked S (safe) and F (fire). The exposed letter shows the
operating state of the weapon.
4.15.1.3 Cocking Handle. The handle (Figure 4-12), at
the right front of the receiver, is used to manually cock the
weapon.
4.15.1.4 Cover Latch. The latch (Figure 4-12) is at the
right rear side of the cover assembly. When the latch lever

4-26

Change 10

is vertical it locks the cover in the closed position. When
moved to the horizontal it unlocks the cover.
4.15.1.5 Grip and Trigger Assembly. The assembly
(Figure 4-12) at the rear section of the receiver, includes
the spade grips. The U-shaped trigger design permits the
weapon to be fired by thumb pressure from either hand.
Table 4-2. Authorized Ammunition

7.62mm:

NATO M59, Ball

7.62mm:

NATO M61, Armor pierce

7.62mm:

NATO M62, Tracer

7.62mm:

NATO M63, Dummy

7.62mm:

NATO M80, Ball

4.15.1.6 Magazine Release. The magazine release
latch (Figure 4-12) is on the left side of the receiver. The
latch spring automatically locks when the ammunition box
is seated on the magazine bracket. Pressing the release latch
manually releases the ammunition box.
4.15.2 Installation of Machinegun M60D.

CAUTION

The RS 890224 pintle mount stop must be
installed in its proper position if the external stores support system (ESSS) is installed. The stop is a three-position stop:
stow, wings only, and external tanks. The
stow and wings only positions are independent of aircraft side. The external
tanks position is particular to the aircraft
side. Care must be taken to ensure the
correct position and/or side is installed.
Use of the M60D machineguns is prohibited when external ERFS tanks are installed on the inboard vertical pylon.
1. Install one machinegun M60D on the right side
and one on the left side of the helicopter at the
crew chief/gunner’s stations.

TM 1-1520-237-10

2. Attach gun to pintle with quick-release pin, and
safety by passing a plastic tie or 0.032-in. safety

Change 8

4-26.1/(4-26.2 Blank)

TM 1-1520-237-10

M60 DEPLOYED POSITION
(SAME BOTH SIDES)

AMMUNITION
BOX

M60 STOWED
(SAME BOTH SIDES)

AMMUNITION AND
GRENADE STORAGE
(SAME BOTH SIDES)

160O

A
85O

75O

FORWARD

AFT

FORWARD

AFT

EJECTOR
CONTROL BAG

1.5o

85O

75O

160O
70O

70O

ELEVATION
AND
DEPRESSION

AZIMUTH

FIELD OF FIRE

AA0411C
SA

Figure 4-11. Machinegun M60D Installation
4-27

TM 1-1520-237-10

BARREL
LOCK
LEVER

AMMUNITION
EJECTION
PORT

FRONT
SIGHT

COVER
LATCH

COCKING
HANDLE

SAFETY

S
GAS
CYLINDER
EXTENSION

AMMUNITION
FEED TRAY
CARRYING
HANDLE

REAR
SIGHT

GAS CYLINDER

BIPOD

QUICK−RELEASE
PIN

RIGHT SIDE

GRIP AND
TRIGGER
ASSEMBLY

SAFETY IN SAFE (S) POSITION

MAGAZINE
RELEASE
LATCH
AA0508
SA

Figure 4-12. Machinegun 7.62 Millimeter M60D
wire through quick-release ring and around the
pintle (Figure 4-14).
F

3. Check that each gun moves freely in azimuth
and can be depressed.
LEFT SIDE

4. Removal of gun is reverse of installation.

SAFETY IS IN
FIRING (F) POSITION
AA0507

4.15.3 Installation of Ejector Control Bag.
1. Position bag on right side of gun (Figure 4-15).
2. Position forward arm bracket of bag in front of
matching forward mounting point on gun
adapter. At the same time, press down on rear
bracket safety latch lever. Slide bag basket
rearward on mounting points and plate.
3. Position rear bracket of bag behind mounting
plate on bottom of receiver.
4. Release latch to lock bag in place.
5. Check bag for positive attachment to gun.

4-28

Figure 4-13. Location and Identification of
Safety on Machinegun M60D
6. Removal of ejector control bag is reverse of
installation.
4.15.4 Installation of Ammunition Can.
1. Open release latch and install can assembly
(Figure 4-16).
2. Check that latch makes positive lock, holding
can in place.

TM 1-1520-237-10

TO INSTALL: POSITION MACHINE
GUN MOUNTING BRACKET IN
PINTLE AND SECURE WITH
MOUNTING PIN
MOUNTING PIN CABLE

REAR
MOUNTING
POINTS
EJECTION
CONTROL
BAG
GUN ADAPTER

QUICK−RELEASE PIN

PINTLE
PLASTIC TIE
AA0506
SA

Figure 4-14. Installation of Machinegun M60D
on Pintle
3. Removal of ammunition can is reverse of installation.

GUN
ADAPTER
FORWARD
MOUNTING
POINTS

FORWARD
ARM BRACKET

STEP 1
POSITION EJECTION CONTROL
BAG ON MOUNTING POINTS

4.15.5 Loading Ammunition.
MOUNTING
PLATE
RECEIVER

WARNING
Observe all safety precautions for uploading ammunition in accordance with TM
9-1095-206-13.
1. With ammunition can installed, retract bolt
fully.

LOCK

REAR BRACKET
SAFETY LATCH

UNLOCKED

STEP 2
REAR BRACKET SAFETY LATCH
LOCKED POSITION

3. Open latch and raise cover assembly.
4. Insert link belt with open side of links down on
tray assembly (Figure 4-17).

AA0505
SA

2. Press safety button to (S) position.

Figure 4-15. Installation of Ejector Control Bag
on Machinegun M60D
live round is in chamber. Move cocking
handle forward by hand.

5. Close cover and latch in place.
4.15.6 Cocking Machinegun M60D.

1. Open ejector control bag and pull cocking
handle fully to rear (Figure 4-18).

WARNING

To prevent accidental firing, do not retract bolt and allow it to go forward if
belted ammunition is in feed tray, or a

Cocking handle assembly must be returned to full forward (locked) before firing.

4-29

TM 1-1520-237-10

RELEASE LATCH
(PRESS TO OPEN)

MAGAZINE
BRACKET

INS

TAL MOVE
L−R
EM
O

LINK BELT

AMMUNITION
CAN ASSEMBLY
AA0693

AA0503

Figure 4-16. Installation and Removal of
Ammunition Can on Machinegun M60D

Figure 4-17. Positioning Cartridge Link Belt on
Machinegun M60D

2. Move cocking handle full forward to locked
position.
3. Press safety button to (S) position (Figure
4-13).

CAUTION

Do not fire machinegun unless the ejector
control bag is mounted in place.
4. With machinegun M60D positioned, loaded,
and aimed, push safety button (Figure 4-13) to
firing (F) position.

4-30

AA0504
SA

Figure 4-18. Charging (Cocking) Machinegun
M60D

5. To fire gun automatically, press trigger fully
and hold. See Figure 4-11 for field of fire.

NOTE

6. Low cycle rate of fire of machinegun M60D
allows firing of single rounds or short bursts.
Trigger must be completely released for each
shot.

When ammunition is exhausted, the last link
will remain in tray assembly. The link assembly can be removed by hand after the
cover assembly is opened for loading.

TM 1-1520-237-10

4.15.7 Firing Malfunctions.

WARNING
If a stoppage occurs, never retract bolt
assembly and allow it to go forward again
without inspecting chamber to see if it is
clear. Such an action strips another cartridge from belt. If an unfired cartridge
remains in the chamber, a second cartridge can fire the first and cause injury to
personnel and/or weapon damage.
a. Misfire. This is a complete failure to fire. It must be
treated as a hangfire until this possibility is eliminated.
b. Hangfire. This is a delay in functioning of the propelling charge. Time intervals set out in paragraph 4.15.8 must
be observed after a failure to fire.
c. Cookoff. This is firing of the chambered cartridge
from a hot barrel. A cookoff may occur from 10 seconds to
5 minutes after cartridge has been in contact with barrel.
d. Runaway Gun. If gun continues to fire after trigger
has been released, grasp belt, twist and break belt, allowing
the gun to run out of ammunition (usually when the belt is
broken only 3 to 5 rounds will remain).
4.15.8 Failure to Fire Procedure.
1. If a stoppage occurs, wait 5 seconds. Pull
handle assembly to rear, making sure operating
rod assembly is held back.
2. If cartridge ejects, return handle assembly forward, re-aim machinegun, and attempt to fire.
If machinegun does not fire, it must be cleared.
3. If cartridge does not eject, retract bolt assembly. Move safety button to SAFE (S), position.
Remove ammunition and links and inspect receiver assembly. Move safety button to FIRING (F) and attempt to fire. If cartridge does
not fire and barrel is considered hot enough to
cause a cookoff (200 rounds fired within 2 minutes), wait 15 minutes with bolt assembly in
forward position. Remove cartridge and reload.
If weapon is not hot enough to cause a cookoff,
disregard 15-minute wait.
4.15.9 Extracting a Ruptured Cartridge Case. Position ruptured cartridge case extractor in chamber. Run

cleaning rod through barrel assembly from muzzle end. Tap
cleaning rod against extractor until extractor and ruptured
cartridge case come out of chamber.
4.15.10 Double Feeding. When a stoppage occurs with
bolt assembly in forward position, assume there is an unfired cartridge in chamber. Treat this as a hangfire (paragraph 4.15.7).
4.15.11 Unloading. Raise cover assembly and remove
linked cartridges. Inspect chamber.
4.15.12 Ammunition. Ammunition for the machinegun
is connected to form a link belt; the rounds are used to hold
two links together. When a round is fired, the cartridge and
link separate and is contained in the ejector bag assembly.
Ammunition stowage in the cabin has compartments for six
grenades and extra rounds of ammunition.
4.16 VOLCANO
SYSTEM. VOL

MULTIPLE

MINE

DELIVERY

The volcano system is an automated, scatterable mine
delivery system that is capable of launching mines from the
helicopter. The system can dispense mines during day/night
and all weather conditions. The system can lay a mine field
of up to 960 mines at an average density of 0.9 mines per
meter front. For a more detailed description of the volcano
system, refer to TM 9-1095-208-23-2&P.
NOTE
The forward two-thirds of the cabin entry/
exit doors are restricted by the volcano system making the loading and unloading of
passengers and cargo more difficult. Internal
loads should be planned accordingly.
4.16.1 System Components. The volcano system consists of five major components: dispenser control unit
(DCU), launcher racks, jettison system, aircraft mounting
hardware, interface control panel (ICP) and ammunition
mine canisters.
4.16.1.1 DCU. The DCU (Figure 4-19) is the primary
electronic component and houses the electronics that control the system and contains a control panel for operating
the system. An interface control panel (ICP) is provided to
connect the applicable controls to the DCU. The DCU controls firing signals to the canisters and conducts built-in
testing (BIT) of the entire mine dispensing system. The
DCU is the main operator interface for the system. On the
DCU TOP panel, the operator controls the system with the
following controls:

Change 9

4-31

TM 1-1520-237-10

a. POWER switch.
b. SELF DESTRUCT TIME control.
c. HELICOPTER DELIVERY SPEED control.
d. FIRE CIRCUIT ENABLE switch.

CONTROL/
INDICATOR
VOLCANO ARM
switch

Arms system.

ARMED light

Illuminates when VOLCANO
ARM switch is pulled out and
moved up if canisters are loaded
and system is enabled. Light will
go out when switch is pulled out
and moved down.

Flashing
light

ARMED

Indicates arm enable or launch
switch error (GA) or all canisters
have been dispensed.

JETTISON TEST/
ARMED
switch
indicator

Press to start/stop jettison system
self-test.

e. DIM control.
f. FAILURE INDEX switch.
g. TEST/OVERRIDE switch.
From the DCU panel the operator also oversees system
status with the SYSTEM FAILURE IDENTIFICATION/
TEST displays. Power to operate the volcano system is
provided from No. 1 dc primary bus through the CMD
CSL SET circuit breaker.
4.16.1.2 Launcher Rack Jettison System. The jettison system consists of an interface panel, emergency and
primary jettison circuits, and explosive cartridges. Jettison
capability is provided for upper and lower launcher racks
on each side of the helicopter. The one shot jettison circuits
are independent of each other and either may be used to
jettison the launcher racks. Upon activation of JETTISON
or EMER JETTISON, the lower racks on each side will
separate from the helicopter. After a lower racks away signal is received by the upper racks, the upper racks will
automatically jettison. The EMER JETTISON is a backup
for the JETTISON and provides the same launcher rack
jettison function on an independent circuit. When the helicopter is on the ground, the weight-on-wheel switch disables the jettison circuits to prevent unintentional jettisoning of the racks. Power for the jettison system is provided
from the dc essential bus through the ESSS JTSN OUTBD
and ESSS JTSN INBD circuit breakers. Power for emergency jettison is supplied by the battery.
4.16.1.3 Interface Control Panel (ICP). Controls and
indicators for the ICP (Figure 4-19) are on the control panel
and are as follows:

CONTROL/
INDICATOR
EMER
JETTISON/
JETTISON switch

4-32

FUNCTION
Jettison all launcher racks with
canisters.

FUNCTION

Illuminates if jettison system passes
test.

Illuminates if jettison cartridges are
missing.

Flashing F

Indicates one of the following
malfunctions:
-Weight-on-wheels switch.
-EMER JETTISON switch.
-JETTISON switch.
-Missing rack sensor(s).

NOTE
VOLCANO ARM function is independent
of EMER JETTISON and JETTISON
function.
4.16.2 Volcano Operational Check.

CAUTION

This operation will be completed only
when helicopter is on the ground.

TM 1-1520-237-10

B
A

EMER
JETTISON

JETTISON TEST
P

JETTISON

ARMED

A
VOLCANO ARM

INTERFACE CONTROL PANEL (ICP)

C
B

SYSTEM FAILURE IDENTIFICATION / TEST
RACK

COLUMN

ROW

QUANTITY

ERROR

TR
IM
ICK
FWD

FAILURE
INDEX

TEST

OVERRIDE

AFT

O
RG
CA EL.
R

CANISTERS REMAINING
POWER

DCU
FAILURE

SET
TIME

LEFT

RIGHT
BRIGHT

SELF DESTRUCT
TIME
1
POWER

RESET

2
3

HELICOPTER
DELIVERY SPEED
102
74
40 55
148
56
30
80
37
27
KPH
KTS

FIRE CIRCUIT
ENABLE

MINE
LAUNCH
CONTROL

222
120
KPH
KTS

OFF
OFF

CYCLIC STICK GRIP
DISPENSER CONTROL PANEL (DCU)
(ON HELICOPTERS EQUIPPED WITH M−139 MINE DISPENSER KIT)

AA9411B
SA

Figure 4-19. Volcano Mine Dispenser Controls VOL

4-33

TM 1-1520-237-10

NOTE
If, for any reason prior to a mission, DCU is
turned OFF or loses power, redo steps 2
through 5.
DCU must warm up for two minutes prior to
mine laying operation.
DCU will not function if backup hydraulic
pump is in operation.

cator, otherwise, mines may be launched
with a false self-destruct time.
5. SELF-DESTRUCT TIME indicator flashes
after 15 seconds. Set self-destruct time by turning SELF-DESTRUCT switch to RESET position for minimum of two seconds, and then to
the desired 1, 2 or 3 setting. The set time display must agree with SELF-DESTRUCT
switch position. If not, repeat procedure. If
wrong indication appears again, postpone mission and return to maintenance.

1. Install mine canisters (TM 1-1520-237-23).
2. Toggle DCU POWER switch to POWER. BIT
runs automatically. POWER indicator displays
ON. Other indicators displays 8s or 88s except
DCU FAILURE display, which will flash F.
All displays, except for POWER are then blank
(Figure 4-20).
3. Displays automatically go blank followed by
88s being displayed in left and right CANISTERS REMAINING displays. If an error code
appears in ERROR display, refer to TM
9-1095-208-23-2&P.

6. Repeat step 5. to reset or change the selfdestruct times.
NOTE
If flashing 1 continues to appear in DCU indicator, repeat setting self-destruct time.
7. Set planned dispensing helicopter ground speed
with HELICOPTER DELIVERY SPEED
knob on DCU.
4.16.3 Arming Canisters.

NOTE

WARNING
The following assumes a full canister load.
Any load other than 80 canisters each side
should result in the number of canisters
loaded being displayed instead of 80 after
overriding the applicable number of error
code 99’s9 for empty canister slots.
4. Toggle TEST/OVERRIDE switch to TEST.
Canister test is initiated. Canister test is complete when 80s are displayed in CANISTERS
REMAINING readout and no error code appears. If an error code appears in ERROR display, refer to TM 9-1095-208-23-2&P.
NOTE
Error codes 5, 8 and 9 may be overridden.
Refer to TM 9-1095-208-23-2&P.

WARNING
Check SET TIME display to make sure
number set with dial is displayed in indi-

4-34

Change 1

Do not walk or stand in front of launcher
racks when racks are loaded and arming
levers are locked in armed position.
NOTE
Green latching levers must be in locked position before red arming levers are moved to
armed position.
Verify red arming levers are fully forward to
the arm (lock) position by pushing back levers without depressing plungers.
Mechanically arm one row of canisters at a time as follows:
1. Individually seat each canister of the five canisters in a row by pushing canisters in and up
into rack keyholes. Canisters should be loaded
top to bottom back to front.
2. Depress plunger on red arming lever with
thumb and pull lever towards personnel. When

TM 1-1520-237-10

SYSTEM FAILURE IDENTIFICATION / TEST
RACK

COLUMN

ROW

DCU
POWER FAILURE

SET
TIME

QUANTITY

ERROR

SYSTEM FAILURE IDENTIFICATION / TEST
RACK

COLUMN

ROW

DCU
POWER FAILURE

SET
TIME

CANISTERS REMAINING
LEFT

QUANTITY

ERROR

CANISTERS REMAINING

RIGHT

LEFT

RIGHT

SYSTEM FAILURE IDENTIFICATION / TEST
RACK

COLUMN

ROW

DCU
POWER FAILURE

SET
TIME

QUANTITY

ERROR

CANISTERS REMAINING
LEFT

RIGHT

(SEE NOTE)

NOTE
READOUT IF NO CANISTERS ARE
INSTALLED.
AA9412
SA

Figure 4-20. Volcano System DCU Displays VOL

4-35

TM 1-1520-237-10

lever begins to move, release thumb from
plunger. When lever reaches locked position
(about 45°) plunger will click out. Push inboard
to ensure lever has locked (Figure 4-21).
3. Repeat steps 1 and 2 for all racks.
4. Remove all eight jettison system REMOVE
BEFORE FLIGHT safety pins from each side
panel.
NOTE
When operating the helicopter configured
with the volcano system at high gross weight
and high airspeeds, the pilot may encounter
intermittent lateral tail pulse 9tail shake9. The
intensity of the tail shake is further aggravated by left sideslip.
5. Lift off and proceed to designated target.
6. Prior to reaching target area remove safety pin
and place the DCU VOLCANO ARM switch
on the ICP in ARMED position (Figure 4-22).
(Pull VOLCANO ARM switch out and lift up.)
7. FIRE CIRCUIT switch to ENABLE.
4.16.4 Mine Launch Control. Launching of mines is
controlled by buttons on the pilot’s cyclic control, marked
GA (go around) (Figure 4-19). When the system is in an
armed condition, pressing and releasing either GA button
will start the launch sequence. If mines are being launched,
press and release either GA button to stop mine launching.
In addition to pressing the GA button to stop mine launch,
the VOLCANO ARM switch may be moved to off (down)
position or the DCU FIRE CIRCUIT switch may be
placed OFF.
NOTE
If launching is stopped by pressing a GA
button before all mines are launched, the
launch must be restarted within 60 seconds
of stopping to prevent an error code. The
VOLCANO ARM switch should also be
moved to off (down) position within 60 seconds of stopping to prevent an error code.

4-36

Change 10

CAUTION

Ensure same number of racks are installed on either side of aircraft.
Ensure same number of canisters are installed on either side of aircraft.
Partial load of canisters may result due to
rack or canister failures (error codes 5, 8
or 9).
Partial canister loads, if not adjusted or
balanced, will degrade mine field pattern.
1. Place TEST/OVERRIDE switch to OVERRIDE.
2. After overriding error code 9, remove and discard all failed canisters.
3. One at a time, fill vacant positions left by removal of failed canisters as follows:
a. Remove canister from top rack of same
side, forward most column, top most position with canister.
b. Place this canister in vacant position.
c. Repeat steps a. and b. above until all positions, other than those removed to fill vacant positions, have canisters.
4. Canister load may be balanced out by moving
canisters from side with most canisters to side
with fewest canisters as follows:
a. Remove canister from top rack, forward
most column, topmost row of side with
canisters.
b. Place removed canister in top rack, rear
most column, lowest position of side with
fewest canisters.

TM 1-1520-237-10

LAUNCHER RACK

DCU

AA9413
SA

Figure 4-21. Arming Volcano System Canisters VOL
c. Repeat steps a. and b. above until load is
balanced.

4. Verify that the HELICOPTER DELIVERY
SPEED settings agree with the helicopter
ground speed.

4.16.5 Mine Launch.
NOTE
Mine launching may be started and stopped
as many times as required until all mines
have been launched.
If launching is stopped by pressing a GA
button before all mines are launched, the
launch must be restarted within 60 seconds
after stopping to prevent an error code. The
VOLCANO ARM switch may also be
moved to off (down) position within 60 seconds after stopping to prevent an error code.
1. DCU FIRE CIRCUIT switch safety pin and
streamer (Figure 4-21) - Remove.
2. DCU FIRE CIRCUIT switch - ENABLE.
3. Before reaching target, VOLCANO ARM
switch (Figure 4-22) - ARM. Verify P/F/
ARMED indicates ARMED.

CAUTION

FIRE CIRCUIT switch must be enabled
for at least two minutes prior to mine
launching.
5. GA button - Press either pilot’s to start launching mines. Press either GA button a second
time to stop mine launching.
NOTE
If launching is interrupted for longer than 60
seconds, the ARMED light on the ICP will
flash and error code 1 will be displayed in
the DCU ERROR indicator.
6. If launch is interrupted longer than 60 seconds,
resume launch: VOLCANO ARM switch - Off
for at least sixteen seconds.

Change 10

4-37

TM 1-1520-237-10

EMER
JETTISON

JETTISON TEST
P

JETTISON

ARMED

VOLCANO ARM

AA9414
SA

Figure 4-22. Volcano ICP VOL
7. VOLCANO ARM switch - ARM. Verify a
steady ARMED is displayed on the ICP.
Launching can then be resumed.
8. During mine launching, if an error code appears
on DCU panel that effects mission performance, do the following:
a. DCU FIRE CIRCUIT switch - OFF.
b. Safety pin and streamer assembly - Install
to FIRE CIRCUIT switch.
c. DCU POWER switch - OFF.
d. Return to down loading area and remove
canisters, refer to TM 9-1095-208-23-2&P.
4.16.6 Volcano System Recycle Procedures. These
procedures are to be used in the event of a volcano system
lock-up during a tactical mission.
NOTE

1. Identify error code.
2. Return the VOLCANO ARM switch on the
ICP to the safe position.
3. Return the FIRE CIRCUIT ENABLE switch
on the DCU to the OFF position.
4. Toggle the POWER switch on the DCU to
OFF, then release. Ensure DCU turns off.
5. Leave the DCU turned off at least two minutes
prior to system restart.
NOTE
During cold temperature operation (less than
0°C (32°F)) , the DCU should be turned off
at least five minutes prior to system restart.
6. After waiting the minimum time, toggle
POWER switch to on, then release. Ensure that
the DCU displays the following sequence:
a. Error code F.

If any error codes occur during this recycle,
return to step 1 and repeat all steps.
4-38

Change 10

b. 9ON 9 displayed under POWER.

TM 1-1520-237-10

c. All 98’s9 in the remainder of the displays.
d. The DCU is blank, except for the 9ON9
message.
7. Wait approximately one minute. At that time,
the DCU should display 988 889 under CANISTERS REMAINING.
8. Toggle the TEST/OVERRIDE switch to
TEST, then release.
9. After approximately two minutes, the DCU
should display 980 809 or what was displayed
upon canister installation under CANISTER
REMAINING.
10. After approximately 15 seconds, the number
under SET TIME should flash. At that time,
turn SELF DESTRUCT TIME knob to RESET for 20 seconds.
11. Turn SELF DESTRUCT knob to the desired
setting. Ensure that this setting is displayed under SET TIME without any flashing.
12. Move the FIRE CIRCUIT ENABLE switch
on DCU to ENABLE.

13. Move the VOLCANO ARM switch on the ICP
to ARM, if the switch was in that position before restarting.
4.16.7 Partial Load Error Codes.
a. Three error codes can be overridden to allow mine
laying without a full load of canisters. These are:
(1) Error code 5 - Rack problems. This allows operation with less than 4 racks, if desired, for laying an abbreviated minefield.
(2) Error code 8 - Rack electronics error. This error
indicates one complete row of ten canisters is not available.
No other errors shall be overridden with the error code 8
override.
(3) Error code 9 - Canister failure. Canister failures
are allowed to be overridden. Failed canister should be removed and remaining canisters balanced prior to mission.
b. Troubleshoot all error codes overridden during mission after completion of flight and make an

Change 1

4-39

TM 1-1520-237-10

appropriate entry on DA Form 2408-13-1, refer
to TM 9-1095-208-23-2&P.
4.16.8 Volcano Post Mission Procedures.

erator will be required to perform additional checks when
operating during extreme climatic conditions.
4.16.9.1 Operating In Extreme Cold.

4.16.8.1 Post Mine Launch Check.
1. ICP VOLCANO ARM switch - Off (down).

CAUTION

2. DCU FIRE CIRCUIT switch - OFF.
3. Safety pin and streamer assembly - Install to
FIRE CIRCUIT switch.
4. DCU POWER switch - OFF.
4.16.8.2 After Landing Checks.

WARNING
Do not stand in front of rack loaded with
mine canisters. All personnel will remain
clear of the outboard side of the launcher
racks until the arming levers are safe and
the jettison safety pins are installed, or
until the helicopter is shutdown and
power removed.
1. Install jettison safety pins (four on each rack).
2. Push in plunger on red arming lever and push
lever back until it is in safe position, parallel
with rack and plunger clicks out. Latch all arming levers in safe position.

Static electricity discharges may damage
DCU.
NOTE
Operators wearing arctic gloves should have
no difficulty installing or operating the volcano system.
1. Check movement of launcher rack arming and
latching levers to ensure that they do not bind.
Use warm air source to loosen if required.
2. Install launcher rack cover when dispenser is
installed, but canisters have not been loaded.
3. Perform complete volcano operational check
before any mission (paragraph 4.16.2).
4. Assure all expando and/or single acting pins
have seated and spring button is out.

1. Remove canisters.

5. Check to see that launcher rack levers and canister connectors are free of ice, snow and frost.
Use warm air source to clean and dry as necessary.

2. Record all error codes overridden during mission on DA Form 2408-13-1.

6. Allow 10 minutes of additional warm up time
before using system.

3. Install launcher rack covers.

7. Prior to turning on power, make sure all DCU
switches are free of ice, that FIRE CIRCUIT
switch and streamer are free of ice, and that
rotary switches move freely. Use warm air to
heat and dry sticking or stiff switches.

4.16.8.3 Post Flight Checks.

4.16.9 Volcano Operation Under Unusual Conditions. The volcano system is designed to operate during
adverse weather and extreme temperature conditions. Op-

4-40

Do not force launcher rack levers or
mounting pins to operate.

Change 10

TM 1-1520-237-10

4.16.9.2 Operating In Wet, Mud, Salt Water and Ice
Conditions.

5. Check launcher rack canister connectors for ice.
Use warm air source to melt and dry connectors.

WARNING

6. Check launcher rack arming and latching levers
for ice. Use warm air source to remove ice.
Test operation of levers to assure they have free
movement.

Wet and/or icy hardware may be slippery.
Use extra precaution when handling dispenser components. Do not force ice covered launcher rack levers.
1. During flight in icing conditions shed ice particles may cause foreign object damage (FOD)
to the helicopter, especially main rotor and tail
rotor blades and engine compressors. Flight
tests have shown that this FOD is difficult to
detect during flight. Minimizing descent rates
after ice has accumulated on the helicopter or
external stores should reduce the probability of
FOD because the airflow will carry particles aft
and down away from the helicopter. Normal
instrument procedure descents or approximately 100 feet per minute (fpm) or less are
preferable. During shutdown, crewmember’s
should be alert for unusual engine noise (high
pitched whine) that indicates compressor damage. The helicopter should be visually inspected
prior to further flight.
2. Engine torque increase of up to 20 percent can
be expected during cruise flights in icing conditions with the volcano system installed.
3. After flight in icing conditions with the volcano
system installed, the jettison safety pins may be
difficult to install due to ice in and around the
safety pin holes. The forward launcher rack
locking levers and arming levers may also be
covered with ice making it difficult to move the
arming levers to the safe position. Use an external heater to remove ice from these areas. Do
not use foreign objects to break ice from these
areas as this may cause damage to the system.
4. Do not bend ice covered cables until ice has
been removed with a warm air source.

7. Check expando and single acting pins to ensure
that they are seated and spring button is out.
8. After exposure to mud or salt water, clean and
wash dispenser components immediately. If
dirty, clean, wash and dry components before
repackaging them into shipping containers, refer to TM 9-1095-208-23-2&P.
9. When DCU cover is removed, make sure that
switches are free of ice. Remove ice using
warm air source.
10. Check DCU connectors for ice. Remove ice using warm air source.
11. Check DCU switch for ice. Remove ice with
warm air source.
12. Allow 10 minutes of additional warm up time
before testing and operating dispenser.
13. Perform PMCS, refer to TM 9-1095-208-232&P.
14. Perform complete volcano operational check
before any mission (paragraph 4.16.2).
4.16.10 Accident Procedures.

WARNING
After an accident, turn DCU power OFF,
evacuate all personnel to a distance of
2000 feet (640 meters) and notify EOD.

4-41

TM 1-1520-237-10

Section III CARGO HANDLING SYSTEMS
4.17 CARGO HOOK SYSTEM.
The system consists of a hook assembly (Figure 4-23)
mounted on the lower fuselage, a control panel on the upper console (Figure 2-7), a normal release button on each
cyclic stick grip, one emergency release switch on each
collective stick grip, and a firing key in the cabin for use by
the crew chief. The system incorporates three modes of
load release, an electrical circuit actuated from the cockpit,
a manual release worked by the crewmember through a
covered hatch in the cabin floor, and an emergency release
system using an electrically-activated explosive charge.
4.17.1 Cargo Hook Stowage. The cargo hook shall be
maintained in the stowed position during periods of nonuse. The cargo hook can be placed in a stowed position
(Figure 4-23) by opening the cargo hook access cover in
the cabin floor, and pulling the hook to the right and up.
When the hook is in the stowed position, the load beam
rests on a spring-loaded latch assembly and is prevented
from vibrating by a teflon bumper applying downward
pressure on the load beam. To release the hook from its
stowed position, downward pressure is placed on the latch
assembly lever, retracting the latch from beneath the load
beam, allowing the cargo hook to swing into operating position.
4.17.2 Cargo Hook Control Panel. The CARGO
HOOK control panel (Figure 4-23), on the upper console,
consists of an EMERG REL NORM, OPEN, SHORT
test switch, a TEST light, CONTR CKPT or ALL station
selector switch and an ARMING, SAFE, ARMED switch.
Before the normal release (electrical) can operate, the
ARMING switch must be at ARMED to provide electrical
power to the release switches. The pilot and copilot
CARGO REL switches, on the cyclics, will release the
load when the CONTR switch is at CKPT or ALL. The
crewmember’s NORMAL RLSE switch will release the
load when the CONTR switch is at ALL. The EMERG
REL switch and TEST light permits checking the emergency release circuit when at SHORT or OPEN. In both
cases of testing, if the release circuit is good, the TEST
light will go on when the HOOK EMER REL switch on
pilot’s or copilot’s collective, or the EMER RLSE switch
on the crewmember’s pendant, is pressed.
4.17.3 Crewmember’s Cargo Hook Control Pendant. The crewmember’s cargo hook control pendant
(Figure 4-24), in the aft cabin, provides the crew chief with

4-42

Change 9

an electrical release and jettison of an external load when
the CARGO HOOK CONTR switch is placed to ALL.
The NORMAL RLSE and EMER RLSE switches are
covered by guards to prevent accidental activation. When
the cover is raised the switch can be pressed. When not in
use, the pendant is stowed in the stowage bag at the back of
the pilot’s seat. Electrical power to operate the pendant is
provided from the No. 2 dc primary bus through a circuit
breaker, marked CARGO HOOK CONTR.
4.17.3.1 Normal Release. Normal release of external
cargo is done by pressing the CARGO REL switch on
either cyclic stick grip or the CARGO HOOK NORMAL
RLSE on the crewmember’s cargo hook pendant, after
placing the CARGO HOOK ARMING switch to
ARMED. A light on the advisory panel will go on, indicating HOOK ARMED. This informs the pilot that electrical power is applied to the control circuit; the actuation
of any of the release switches will release the load. When
the CARGO REL switch is pressed and the release solenoid begins to move, a switch closes, lighting the CARGO
HOOK OPEN advisory light. The load arm will swing
open, releasing the cargo. When the sling is detached from
the load beam, spring tension on the arm will cause it to
close and relatch, putting out the CARGO HOOK OPEN
advisory light. The normal release system is a one-shot
cycle; once the solenoid travel begins and the load arm
relatches, the release cycle can again be initiated. Power to
operate the normal release system is supplied from the No.
2 dc primary bus through circuit breakers marked CARGO
HOOK CONTR and PWR.
4.17.3.2 Operational Check - Normal Release
Mode.
1. CARGO HOOK CONTR switch - As required. CKPT for pilot and copilot check, or
ALL for crewmember check.
2. CARGO HOOK
ARMED.

ARMING

switch

3. HOOK ARMED advisory light - Check on.
4. Place about 20 pounds downward pressure on
load beam.

TM 1-1520-237-10

5. CARGO REL switch (pilot and copilot);
NORMAL RLSE (crewmember) - Press and
release.

6. Load beam - Check open. CARGO HOOK
OPEN advisory light - On.

Change 10

4-42.1/(4-42.2 Blank)

TM 1-1520-237-10

E
A

EMERGENCY
RELEASE
SWITCH

CARGO HOOK

HOOK
EMER REL

EMERG REL
TEST
NORM
O
P
E
N
SHORT

O
RG
CA EL.
R

CONTR
CKPT

ARMING
SAFE

ALL

ARMED

CARGO HOOK CONTROL PANEL

NORMAL
RELEASE
SWITCH

COLLECTIVE STICK GRIP

CYCLIC STICK GRIP

LATCH
ASSEMBLY
LEVER
BUMPER
CARGO HOOK
ACCESS DOOR

CABIN
FLOOR

CARGO LOAD
BEAM

CARGO HOOK STOWAGE

AA0367_1B
SA

Figure 4-23. Cargo Hook System (Sheet 1 of 2)
4-43

TM 1-1520-237-10

2. Place about 20 pounds downward pressure on
load beam.

E
COVER
(EXPLOSIVE CARTRIDGE
UNDER COVER)

3. Manual release lever - Pull up/turn fully clockwise and release.

MANUAL
RELEASE
LEVER

4. Load beam - Check open.
5. CARGO HOOK OPEN advisory light - On.
6. When downward pressure is released, load
beam will close and latch.

OPEN

STA
343.0

STA
363.0

7. CARGO HOOK OPEN advisory light - Off
when hook closes.

KEEPER
LOAD BEAM

FRONT

CARGO HOOK
AA0367_2A
SA

Figure 4-23. Cargo Hook System
(Sheet 2 of 2)
7. CARGO HOOK OPEN advisory light - Check
off when hook closes.

4.17.4.1 Cargo Hook Emergency Release Circuit
Check.

8. Repeat steps 4. through 7. for copilot and crewmember position.

1. EMERG REL TEST light - Press. Light
should be on.

4.17.3.3 Manual Release. Manual release of external
cargo can be done from the cabin, through a covered port in
the floor, or by ground personnel from outside the helicopter, with power on or off. Turning the release control on the
right side of the hook (Figure 4-23) clockwise, causes the
latching mechanism to release the load beam. The load
beam will not move unless a downward pressure is exerted
to cause opening. With power applied to the helicopter and
the CARGO HOOK ARMING switch at ARMED, the
CARGO HOOK OPEN advisory light will go on at the
start of release control turning, and will go off at the end of
release control rotation.
4.17.3.4 Operational Check - Manual Release
Mode.
1. Manual release lever spring - Installed. Check
that spring is straight and provides positive
pressure on the lever.
4-44

4.17.4 Emergency Release Circuit Tester. The
cargo hook emergency release circuit tester (Figure2-7)
marked CARGO HOOK EMERG REL on the upper console, contains a test light and switch. The test light, marked
TEST, goes on during circuit testing to indicate the system
is functioning properly. The switch, with marked positions
NORM, OPEN, and SHORT, is normally at NORM.
When the switch is placed to OPEN or SHORT and the
cargo HOOK EMER REL switch on the pilot’s or copilot’s collective, or EMER RLSE switch on the crewmember’s cargo hook control pendant is pressed, the circuit
tester light will go on if the circuit is good.

Change 7

NOTE
To prevent unintentional discharge of the
cargo hook explosive cartridge, the pilot
shall call off each procedural step of the
emergency release circuit test before that
step is done. Station being checked shall reply to pilot’s command.
2. Pilot’s release - Check.
a. Short test.
(1) CARGO HOOK EMERG REL
switch - SHORT.
(2) Pilot’s HOOK EMER REL button Press and hold.
(3) CARGO HOOK TEST light - On.

TM 1-1520-237-10

CARGO HOOK
NORMAL RLSE
SWITCH
NORMAL
RELEASE
SWITCH
GUARD

A
CARGO HOOK

STRAP
O
RG
CA O K
HO

PILOT SEAT

L
A
RM
NO
SE
RL

EMERGENCY
RELEASE SWITCH
GUARD

E
RL ME
SE R

EMER
RLSE
SWITCH

STA
343.0

STA
363.0

CREWMEMBER’S CARGO
HOOK PENDANT STOWAGE
RIGHT SIDE

CREWMEMBER’S CARGO HOOK
CONTROL PENDANT
AA0368
SA

Figure 4-24. Crewmember’s Cargo Hook Control Pendant
(4) HOOK EMER REL button - Release.
TEST light off.
(5) Repeat steps (2) through (4) for copilot’s HOOK EMER REL button, and
crewmember’s cargo hook control
pendant EMER RLSE button.
b. Open test.
(1) CARGO HOOK EMERG REL
switch - OPEN.
(2) Pilot’s HOOK EMER REL button Press and hold.
(3) CARGO HOOK TEST light - On.
(4) HOOK EMER REL button - Release.
TEST light off.
(5) Repeat steps (2) through (4) for copilot’s HOOK EMER REL button, and
crewmember’s cargo hook control
pendant EMER RLSE button.
3. CARGO HOOK EMERG REL switch NORM. If the cargo hook is not to be used

immediately after completing the circuit test
check, the EMERG REL switch shall remain
at OPEN until ready for load pickup.
4.17.4.2 Emergency Release.
NOTE
When the emergency hook release has been
used and a replacement squib (explosive cartridge) is not available, the hook can not be
used until the explosive device is replaced,
since the hook load beam will not close and
lock.
Emergency release of an external cargo load is done by
an electrically-fired explosive cartridge, initiated from either of the collective stick grip switches, marked HOOK
EMER REL, or the crewman’s cargo hook control pendant, marked EMER RLSE. The emergency release is used
when the electrical and manual releases are inoperative,
and the load must be jettisoned. With the CARGO HOOK
EMERG REL switch at NORM, power will be applied to
the emergency release switch. Pressing the switch applies
28 vdc to the explosive cartridge, producing a high gas
pressure to drive a piston in the lock assembly, releasing
the load arm lock. The weight of the load will cause the

4-45

TM 1-1520-237-10

load arm to open. Once the emergency release is used, the
hook will remain open and the CARGO HOOK OPEN
advisory light will remain on until the explosive cartridge
device is replaced. When the explosive cartridge device is
replaced the load arm will close, the light will go off, and
the emergency release mode is returned to operation. Power
to operate the emergency release system is from the dc
essential bus through a circuit breaker, marked CARGO
HOOK EMER.
4.17.5 Preflight. When cargo hook loads are to be carried, checks within this paragraph and procedures of paragraphs 4.17.6, 4.17.7, 4.17.8 and 4.17.9 apply.
1. Cargo hook - Check condition, security and explosive cartridge installed.
2. Emergency release system - Check. (Go to
paragraph 4.17.4.)
3. Manual release - Check. (Go to paragraph
4.17.3.3.)
4.17.6 Before Takeoff.
1. CARGO HOOK EMERG REL switch NORM.

4-46

2. CARGO HOOK
ARMED.

ARMING

switch

4.17.7 Emergency Release Procedure.
Pilot or copilot HOOK EMER REL or crewman’s control pendant EMER RLSE - Press.
4.17.8 In-flight Procedures.

CAUTION

Cargo suspended from the cargo hook
should not be over a 30° cone angle. To
prevent damage to the cargo hook keeper,
the pilot shall use extreme care to prevent
placing load pressure on the keeper.
CARGO HOOK ARMING switch - As required.
ARMED for low altitude/low airspeed. SAFE at cruise altitude and airspeed.
4.17.9 Before Landing.
CARGO HOOK
ARMED.

ARMING

switch

TM 1-1520-237-10

Section IV MISSION FLEXIBLE SYSTEMS
4.18 MISSION READINESS CIRCUIT BREAKER
PANEL.
The mission readiness circuit breaker panel (Figure
2-20) is on the No. 1 electrical junction box in the cabin,
and contains all required circuit breakers for mission equipment.

eration in mid-travel. Cable over travel
should not exceed ten feet.

CONTROL/
INDICATOR

FUNCTION

BOOM switch

Swings hoist boom in or out from
cockpit.

4.19 RESCUE HOIST SYSTEM KIT.
The high performance, two speed rescue hoist (Figure
4-25) is post-mounted in the cabin on the right side of the
helicopter when installed. The hoist system consists of
modular components, electrically-driven and electronicallycontrolled, to provide maximum lift capacities of 300
pounds at 0 to 250 feet-per-minute and 600 pounds at 0 to
125 feet-per-minute. The heaviest weight that may be suspended from the hoist is 600 pounds. A speed mode switch
at the back of rescue hoist control panel assembly on the
hoist support assembly, provides a selection of either
SLOW speed (0 to 125 feet-per-minute), or FAST speed (0
to 250 feet-per-minute). The hoist motor mounted at the top
of the pole provides selectable 125 or 250 feet-per-minute
reel-in and reel-out drive of a 250-foot hoist cable. A fail
safe mechanism limits the induced loading to the hoist to
1200 pounds at all times. A continuously running circulating fan cools the hoist motor. The hoist is controlled
through a lower console mounted RESCUE HOIST CONTROL panel and/or crewman’s control pendant grip in the
cabin. A hoist cable cutter system is used to cut the hoist
cable in case of emergency, by exploding a squib-actuated
cable-cutter. The cut cable then drops free of the hoist
boom. Power to operate the rescue hoist system is from the
No. 2 dc primary bus through a circuit breaker on the mission readiness circuit breaker panel, marked RESQ HST
CONTROL. Power for the cable cutter system is from the
dc essential bus through a circuit breaker, marked HOIST
CABLE SHEAR. Refer to Table 2-4 for servicing.
4.19.1 Controls and Function. The RESCUE HOIST
CONTROL panel (Figure 4-25) has all necessary controls
for operating the hoist from the cockpit, and contains the
system MASTER switch, controlling ON or OFF for both
cockpit and cabin. The hoist will respond to the first control
signal received.

OFF

Static position, removes electrical
power from hoist boom positioning
motor.

Provides power to boom motor to
position boom inboard from
cockpit.

OUT

Provides power to boom motor to
position boom outboard from
cockpit.

MASTER switch
OFF

Disconnects all electrical power
from hoist operating controls.

Provides power to both cockpit
controls and crewman’s pendant for
hoist operation.

CABLE

Provide cable up or down control
from cockpit.

OFF

Static position, removes electrical
power from hoist reel motor for
cockpit operation.

Provides power to hoist reel motor
to reel in cable operation from
cockpit at 250 feet-per-minute
only.

DOWN

Provides power to hoist reel motor
to reel out cable operation from
cockpit at 250 feet-per-minute
only.

SQUIB switch

NOTE
TEST
During hoist operation, over travel of the
cable assembly may occur in the extended
mode of operation after stopping hoist op-

Selects control point for hoist
operation.

Selects either TEST or NORM
operation.
Checks condition of CABLE
SHEAR circuit through squib to
indicate circuit is complete.

4-47

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

NORM

Places squib circuit in a ready for
fire condition.

IND

Lights when test of CABLE
SHEAR circuit through squib is
good.

CABLE SHEAR
switch

Controls cable cutter firing circuit.

FIRE

Directs electrical power to cable
cutter squib for shearing hoist
cable.

SAFE

Removes electrical power from
cable cutter circuit.

4.19.2 Boom Assembly Module. The boom assembly
module consists of the boom structure boom head, up limit
switch, cable-cut device, and a cable guide, all installed in
the boom. The boom head is allowed to swivel 65° above
the boom cable and 30° from side-to-side and guide the
cable to wrap or unwrap from a 30° cone angle. The upper
limits of cable control includes an automatic means for decelerating the cable to 67 feet-per-minute cable speed. At 8
to 10 feet below the boom head, a caution light on the
crewman’s control pendant marked CAUTION will go on.
The cable will again decrease speed to 12 feet-per-minute
at 12 to 18 inches below the boom head.
4.19.3 Limit Switches. Four limit switches are actuated
by actuation assembly cams. They are: a down all stop, that
actuates when 250 feet of cable is reeled out; a down limit
switch, that actuates at 247 feet, to provide deceleration; a
10-foot caution switch that actuates when the hook is within
10 feet of the boom head or within 10 feet of the down
limit (240 feet); and an up deceleration switch, that actuates
when the cable hook is within 12 inches of the boom head.
4.19.4 Crewman’s Control Pendant Grip. The crewman’s control pendant grip (Figure 4-25) is a hand-held

4-48

Change 7

control in the cabin. The pendant grip is connected to the
control box by a cable connector. The control pendant contains three switches and two caution/warning lights:
HOIST cable control, BOOM positioning, and ICS
switches; OVERTEMP and cable 10-foot CAUTION
lights. The HOIST control is a directional and variable
speed spring loaded to center switch, with marked positions
of OFF, UP and DOWN. As the switch is moved further
away from OFF, the hoist speed increases in the marked
direction. When the switch is released the hoist will stop.
The BOOM position switch, with marked positions OUT
and IN, operates in the same manner as the HOIST switch,
except the boom moves in or out at a single speed. Two
lights are installed on the crewman’s control pendant. They
are: The 10-foot CAUTION light to warn the crewman
whenever the hoist cable is 10 feet or less from the all stop
limits. A red OVERTEMP light that warns the crewman of
an overtemperature condition in either the hoist lubrication
systems or the hoist motor. Whenever the OVERTEMP
light is on, the hoist should be allowed to cool down until
the light goes off. The ICS control switch, on the front of
the pendant, provides the operator with inter-helicopter
communication.
4.19.5 Cable Shear System. A cable shear feature releases a rescue hoist load in case of an emergency. The
system consists of a squib-actuated cable cutter, a CABLE
SHEAR switch on the pilot’s control panel and a shear
switch at the hoist assembly and a SQUIB test circuit. The
cutter may be fired by the pilot or the copilot from the
SHEAR switch on the control panel (Figure 4-25) or by the
hoist operator using the CABLE-CUT switch on top of the
control box. The SQUIB test circuit consists of a TESTNORM switch and a test good IND light. When the
SQUIB switch is at NORM and the SHEAR switch is
placed to FIRE, electrical power is sent to the dual squib
for firing. The exploding cartridge then drives a cutter into
the hoist cable and shears it. The rescue hoist cable shear
feature is operational whenever power is applied to the helicopter. Once fired, a replacement cable cutter kit and cable
must be replaced. Power to operate the cable shear is provided from the dc essential bus, through a circuit breaker
marked HOIST CABLE SHEAR.

TM 1-1520-237-10

4.19.6 Operation.

hand motion to cable in direction opposite to oscillation.
For significant oscillation, stop reel-in, start reel-out or call
for pilot to lower aircraft.

WARNING
It is the hoist operator’s responsibility to
assure that the hoist cable does not contact any portion of the aircraft. The rescue hoist cable must be kept clear of all
parts of the aircraft and free from other
external obstacles when operating the
hoist. Cable abrasion during hoist operations can lead to cable failure. If cable
contact or snagging occurs, interrupt
hoist operations and inspect the cable for
damage in accordance with applicable
procedures. If any broken wires, unraveling, or kinks are observed, hoisting operations should be discontinued and the cable
replaced.
Reeling a kinked/damaged cable into the
hoist may cause a hoist jam condition
when reel-out is attempted, rendering
hoist inoperative.
The hoist operator is responsible to maintain stability of
the hoisted load by use of hoist controls, ICS calls to pilot,
and physical control of cable (hand or foot). For minor
oscillation (linear or circular swing), stop reel-in, apply

WARNING
If the oscillation is not quickly stopped, it
may become unmanageable. Reeling in an
oscillating load will only aggravate the
motion.
All crew should watch for shock loads, jerks, or snaps
that impart high loads on cable. If observed, hoisting should
be interrupted and cable inspection undertaken to verify
integrity (no broken wires, unraveling, or kinks) before resuming operations. Refer to TM 55-1680-320-23&P, high
performance rescue hoist assembly.
4.19.6.1 Preflight (If use is anticipated).
NOTE
For preflight PMCS, refer to TM 55-1680320-13&P.
1. Check oil level:
a. Release reaction arm and pivot hoist to operating position.

Change 10

4-48.1/(4-48.2 Blank)

TM 1-1520-237-10

MASTER
SWITCH

RESCUE HOIST CONTROL

OUT

FIRE

S
H
E
A
R

SQUIB

CABLE
UP
O
F
F

CABLE SHEAR

MASTER
ON
O
F
F

BOOM
OFF

DOWN

TEST

IND

NORM

SAFE

MISSION−READY CIRCUIT
BREAKER PANEL

C
OIL LEVEL
SIGHT GAGE

HOIST
MOTOR

BOOM

FIL
WI L TO
T
FO H M CEN
−4 Ro O OBILTER
0 P
WI F D ERA ATFOF G
TH RA TI D L
SH IN ON EXRASS
EL AN BE 22
L D D LO 0
ON REF W
AX ILL
T−
1

HELICOPTER
POSITION
SWITCH

UP−LIMIT
SWITCH
LEVER

B
KEEPER

HI
SP GH
EE
D

SPEED
MODE
SWITCH

SQUIB

RESCUE
HOOK

LAPSED
TIME
INDICATOR

F
FR

L
SPOW
EE
D

CONTROL
BOX

HOIST
POST

E
D
10 FOOT
WARNING
LIGHT

CA
OU UTIO
T
N

HO
DO

FIL
WI L TO
T
FO H M IND
OB ICA
−4 R
0o OP IL TE
WI F D ERA ATF D L
TH RA TI D EV
SH IN ON EXR EL
EL AN BE 22
L D D LO 0
ON REF W
AX ILL
T−
1

IST

TE
M

AD
D

BOOM
POSITION
SWITCH

POWER ON
INDICATOR
CABLE
CUT
SWITCH

O
FR

HOIST
WINCH
CONTROL
SWITCH

ICS
SWITCH

AIN
FR

AA0370A
SA

Figure 4-25. Rescue Hoist Kit

Change 5

4-49

TM 1-1520-237-10

b. Check oil level in hoist and boom head.
c. Return hoist to stowed position and secure
reaction arm.
2. Check upper attachment (make sure hose clamp
is installed).
3. Check lower attachment (mounting plates, pip
pins, and star plate).

WARNING
Hands must be kept off hoist boom during
operation to prevent hand entrapment
and injury.
4. Hoist operator - BOOM switch - OUT and then
IN.

4. Check position switch (positions 2 and 4).
5. Ensure hoist main power cable cannon plug is
safetied at junction box.

5. RESCUE HOIST CONTROL panel - Rotate
boom OUT; then IN, then OUT to test boom
operation.

6. Cable cut switches - Down and safetied.

6. Speed mode switch - HIGH.

7. Make sure metallic shorting strip is removed
from cable cut cannon plug.

WARNING

8. Cable cutter connector attached.
9. Check recovery devices are functional and
complete. Make sure recovery devices are secure.
10. Make sure crewmembers have proper personal
equipment (safety harness, leather gloves, and
proper visors).
11. Hoist control circuit breaker - In (mission essential circuit breaker panel).
4.19.6.2 Rescue Hoist Squib Circuit Test.
1. SQUIB switch - Hold at TEST.
2. SQUIB IND light - Check on.
3. SQUIB switch - Release to NORM. SQUIB
IND light off.
4.19.6.3 Boom Position and Hoist Cable Control
Operational Check.
To position the rescue hoist inboard or outboard, do this:
1. MASTER switch - ON.
2. Hoist operator - Check power on indicator (blue
light), check yellow caution light on control
pendant is on, and cooling fan operating.
3. Check ICS switch on pendant.
4-50

Change 10

Rescue hoist cable is stiff and abrasive.
Broken cable strands are sharp, therefore
leather work gloves must be worn whenever handling rescue hoist cables.
A crewmember must reel cable out from
the boom head in line with the boom axis
during the following test procedures. Care
must be taken not to pull the cable taut
around the cable guide/roller, since kinking of the cable might result. Avoid damaging cable on rough surfaces, including
the ground.
7. RESCUE HOIST CONTROL panel DOWN; lower cable hook until all cable is
reeled out. The caution light should extinguish
after the first ten feet of cable is reeled out.
While reeling the cable out, push up on the actuator arm to ensure proper operation of the uplimit switch. Hoist shall stop running when uplimit switch is activated and resume operation
when actuator arm is released.
8. RESCUE HOIST CONTROL panel - Reel in
cable and observe that cable speed slows when
caution light goes on (8 to 10 feet of cable out).
9. Boom up limit actuator arm - Push up on arm
during reeling in to check that hoist stops running when up limit switches are activated. Observe that cable slows when hook is 12 to 18
inches from full up position.

TM 1-1520-237-10

10. Repeat steps 7. through 9., using control pendant assembly. Check that cable speed can be
regulated by control pendant from 0 to 250 fpm
when cable is reeled out beyond 10 feet.
11. Speed mode switch - LOW. Repeat steps 7.
through 9, using control pendant assembly.
Check that cable speed can be regulated by
control pendant from 0 to 125 fpm when cable
is reeled out beyond 10 feet.

CAUTION

Make sure hoist cable is completely up, to
prevent cable wear between cable and
hook assembly.
12. BOOM switch - Rotate boom in to stowed position.
4.19.7 Stopping Procedure.
MASTER switch - As desired.
4.20 AUXILIARY ELECTRICAL CABIN HEATER.
(ON HELICOPTERS EQUIPPED WITH AUXILIARY
CABIN HEATER KIT.)
A 55,000 BTU/hr electrical auxiliary cabin heater is installed in the transition section to provide an increase in
cabin temperature in extremely cold environments. The

auxiliary heater system consists of a heater control panel on
the lower console replacing the retransmission control panel
when the heater kit is installed (Figure 4-26), a blower and
electrical heater unit in the transition section, a heater inlet
port on the cabin aft bulkhead (Figure 2-5), a temperature
control located under the left gunner’s window, and ducting
throughout the cabin. The auxiliary heater system is turned
on from the cockpit by a switch, marked OFF-ON-RESET
on the AUX CABIN HEATER control panel. With both
main generators operating and AUX CABIN HEATER
switch ON, the HTR ON light on the control panel will go
on indicating that power is applied through the heater control relay to the duct temperature sensor and to the blower
motor and cabin heater elements. If a heater unit overheats,
an element thermostat circuit will automatically open, causing the heater to shut off. When the element cools, the
AUX CABIN HEATER switch must be momentarily
placed at RESET to restore the system. A heater outlet
duct cycling thermostat is also provided at the air outlet
side of the heater. If duct temperature exceeds 82° 6 8°C,
the temperature sensor contacts will open, temporarily interrupting power to the heater elements. On decreasing
temperature, the contacts will automatically reset to closed
at 77° 6 8°C to restore power. A redundant duct overheat
sensor/shutoff switch is installed next to the duct cycling
thermostat to shut off power to the heater elements if the
delivered air flow temperature exceeds 99° 6 8°C due to
heater cycling sensor failure. Sensor switch contacts reset
closed on decreasing temperature. However, the cockpit
heater panel switch must be momentarily placed at RESET
to restart system operation. Heated air is carried through the
cabin via ducts along each side of the cabin at the ceiling.

Change 10

4-50.1/(4-50.2 Blank)

TM 1-1520-237-10

ports on the tanks. Control of the system is provided by a
control panel on the lower console. Power to operate the
fuel transfer system is provided from the No. 1 dc primary
bus through circuit breakers marked EXT FUEL LH, and
NO. 1 XFER CONTROL and from the No. 2 dc primary
bus through circuit breakers marked EXT FUEL RH and
NO. 2 XFER CONTROL and from the No. 2 ac primary
bus through a circuit breaker marked AUX FUEL QTY on
the mission readiness circuit breaker panel.

AUX CABIN HEATER
RESET

HTR
ON

HTR
INOP

ON
OFF

(ON HELICOPTERS EQUIPPED WITH AUXILIARY CABIN HEATER)

4.22 EXTERNAL EXTENDED RANGE FUEL SYSTEM KIT. ERFS

4.22.1 External Extended Range Fuel Transfer
Modes. Fuel can be transferred from external tanks to
main tanks in either of two modes, AUTO MODE or
MANUAL. AUTO (primary) transfers fuel automatically
after switches are manipulated. Fuel transfer will be managed by the microprocessor as described in paragraph
4.22.6. The pilot need only occasionally monitor the AUXILIARY FUEL MANAGEMENT panel to ensure that
during AUTO MODE of fuel transfer, fuel in external
tanks is decreasing as it should. The second mode of transfer is the MANUAL XFR (secondary) mode. In the
MANUAL mode the pilot may replenish main tank fuel in
any quantity or frequency desired. Transfer must be initiated by the pilot. The pilot must constantly monitor the fuel
quantity indicator in order to start and terminate transfer to
remain within CG limits. It is possible to transfer fuel from
any one tank while in MANUAL mode. Transfer is shut off
by the pilot when the external tank low-level sensor signals
that the tank is empty. During manual transfer, at the illumination of a tank EMPTY light, immediately switch from
OUTBD to INBD or to manual transfer MODE OFF. Do
not wait for the NO FLOW light to illuminate. This will
preclude air from entering the fuel line and entering the
main tank. At the illumination of the TANK EMPTY capsule, 2.5 to 4.0 gallons of fuel remain in the tank. Sloshing
of the fuel will cause frequent illumination of the TANK
EMPTY, NO FLOW, and AUX FUEL lights when the
tank is in this condition.

The external extended range fuel system is supported by
the external stores support system extending horizontally
from each side of the fuselage aft of the cockpit doors. The
230-gallon and 450-gallon jettisonable tanks may be suspended from the vertical stores pylons (VSP). Removable
fuel lines, bleed-air lines, valves, and electrical connectors
are within the horizontal stores supports (HSS). A tank
pressurizing system, using bleed-air, transfers fuel to the
main tanks. Fuel lines carrying fuel to the No. 1 and No. 2
main fuel tanks contain check valves to prevent backflow.
The extended range system does not supply fuel directly to
the engines but is used to replenish the main tanks. Servicing of the external tanks can be done only through fueling

4.22.2 External Extended Range Fuel System
Tanks. External extended range system contains two or
four tanks suspended from supports outboard of the fuselage. The tanks contain baffles to prevent fuel sloshing.
Quick-disconnect valves are provided in external fuel and
bleed-air lines to provide seals when tanks are jettisoned or
removed. If tanks are not installed cccc will be displayed in
the AUX FUEL QTY POUNDS display when OUTBD or
INBD is selected on the rotary fuel quantity selector. The
preferred location of the External Extended Range Fuel
System (External ERFS) auxiliary fuel tank is the outboard
pylon. This facilitates ingress and egress of troops, loading
of cargo, and the use of the M60D door gun.

LOWER CONSOLE

AA0400
SA

Figure 4-26. Auxiliary Cabin Heater Control
Panel
Power to operate the auxiliary cabin heater elements is provided from the No. 1 ac primary bus through the No. 1
junction box and protected by current limiters. Blower
power is provided from the No. 2 ac primary bus and protected by a circuit breaker, marked AUX HTR BLOWER.
Control of both heater elements and blower is provided by
power from the No. 1 dc primary bus through a circuit
breaker, marked AUX HTR CONTROL. The auxiliary
cabin heater will operate with APU power on and the
backup pump off or with both generators on and the backup
pump off.
4.21 ROTOR BLADE DEICE KIT.
Refer to Chapter 2 for description and operation of the
blade deice kit.

Change 10

4-51

TM 1-1520-237-10

4.22.3 Auxiliary Fuel Management Control Panel.
The AUXILIARY FUEL MANAGEMENT control panel
(Figure 4-27) contains all controls for operating the external extended range fuel system. Controls description is as
follows:
CONTROL/
INDICATOR
FUEL XFR

Controls fuel management of
external extended range system.

OUTBD ON

Opens bleed-air valves to outboard
tanks for pressurization.

OFF

Closes bleed-air valves to tanks.

INBD ON

Opens bleed-air valves to inboard
tanks for pressurization.

OFF

Closes bleed-air valves to tanks.

TANKS INBD

Selects fuel transfer valves from
inboard tanks for fuel transfer to
main tanks; deselects outboard
valves.

TANKS OUTBD

Selects fuel transfer valves from
outboard tanks for fuel transfer to
main tanks; deselects inboard
valves.

AUTO

OFF
MANUAL

4-52

Selects
AUTO-OFF-MANUAL
mode of fuel transfer from external
fuel tanks.
Automatically transfers fuel to
main tanks from selected external
tanks, until empty sensor in each
tank interrupts transfer. Transfer
occurs in levels as shown under
fuel transfer sequence. When tanks
are empty, NO FLOW and
EMPTY indicators and AUX
FUEL caution light will be turned
on.
Interrupts automatic or manual
transfer mode of operation.
Provides
electrical
path
to
MANUAL XFR switch(es), which
allows transfer from selected
tank(s) until switch is moved to off.

Change 8

FUNCTION

MANUAL XFR
RIGHT ON

Transfers from right tank used in
conjunction with MODE switch in
MANUAL position of the pair
selected by TANKS select switch.

OFF

Interrupts transfer operation.

LEFT ON

Transfers from left tank of the pair
selected by TANKS select switch.

OFF

Interrupts transfer operation.

FUNCTION

PRESS

MODE

CONTROL/
INDICATOR

AUX FUEL QTY
POUNDS

Indicates pounds of external fuel
remaining in symmetrical pair of
tanks total of auxiliary tanks, selftest indication or failure codes.
Displays K factors of flow meter.
NOTE

Fuel tanks selector and quantity
indicators are also used in
conjunction with INCR-DECR
switch when initializing fuel
quantity of tanks.
OUTBD

Total pounds of fuel remaining in
outboard pair of tanks.

INBD

Total pounds of fuel remaining in
inboard pair of tanks.

TOTAL

Pounds of fuel remaining in all
external extended range tanks.

CAL

Adjusts K factor of flow switch on
AUXILIARY
FUEL
MANAGEMENT panel.

INCR
position

switch

Increases setting of digital readout
to adjust for fuel remaining in tanks
selected by fuel tank selector.

DECR
position

switch

Decreases setting of digital readout
to adjust for fuel remaining in tanks
selected by fuel tank selector.

STATUS button

Resets AUX FUEL caution light
and stores condition of NO FLOW
and EMPTY indicators.

TM 1-1520-237-10

CONTROL/
INDICATOR
TEST

DEGRADED light

FUNCTION
Checks display and indicator lights.
Performs
memory
checksum,
displays 8 sequentially in each
digital display. Verifies that
temperature probe is connected;
verifies that flow meter is
connected;
performs
trial
calculation based on a known
temperature and flow meter input,
compares it with a known good
value, and displays setting of fuel
density switch. At completion of
test, GOOD or error code will be
displayed (Table 4-3).
Error in critical function has
occurred. Error code will be
displayed as shown under TEST.
Only E02 error will allow
microprocessor to clear failure code
and
regain
fuel
remaining
information by doing two self-tests.

EXTERNAL
*RIGHT NO
FLOW light

Fuel flow does not exit from
selected right tank.

*INBD EMPTY
light

Right inboard tank fuel exhausted.

*OUTBD
EMPTY light

Right outboard tank fuel exhausted.

*LEFT NO
FLOW light

Fuel flow does not exit from
selected left tank.

*INBD EMPTY
light

Left inboard tank fuel exhausted.

*OUTBD
EMPTY
light

Left outboard tank fuel exhausted.

VENT SENSOR
*OVFL

Detects the presence of fuel on the
vent thermistor.

CONTROL/
INDICATOR

FUNCTION

NOTE

Illumination of this capsule on the
Fuel Management Panel will cause
illumination of the AUX FUEL
caution light and the MASTER
CAUTION light. Pushing the
STATUS button will reset the NO
FLOW, AUX FUEL and the
MASTER CAUTION lights, but
does not correct the no flow
condition.
FAIL

Open in vent sensor line.

4.22.4 External Extended Range Fuel Quantity Indicating System. The AUX FUEL QTY POUNDS
digital readout (Figure 4-27) displays the amount of fuel
remaining. Fuel type is preset by switches in the AUXILIARY FUEL MANAGEMENT panel. Preset can only be
done when the helicopter weight is on the wheels. When
measuring quantity, the readout is used in conjunction with
the rotary selector switch to select tank pair subtotal, or
TOTAL remaining in all tanks. Fuel used is sensed from a
common flow transmitter within the fuel line to the main
tanks. This amount is subtracted from the preset fuel quantity input and is displayed on the digital readout as pounds
remaining. A DEGRADED light will go on when a complete failure has occurred in the microprocessor, or an error
condition is detected by the microprocessor, or when the
temperature sensor has failed. Power for the fuel quantity
subsystem is provided from the No. 2 ac primary bus
through a circuit breaker marked AUX FUEL QTY, on the
mission readiness circuit breaker panel.
4.22.5 Auxiliary Fuel Management Control Panel
Test.
1. TEST button - Press. Digits should display 8s,
all panel lights and DEGRADED and VENT
SENSOR (FAIL and OVFL) lights should illuminate.

Indicates fuel venting overboard.

Change 6

4-53

TM 1-1520-237-10

A
AUXILIARY FUEL MANAGEMENT
FUEL

PRESS

OUTBD
ON

INBD
ON

XFR

TANKS
INBD

MANUAL XFR

MODE
AUTO

RIGHT
ON

LEFT
ON

MANUAL

OFF

O
F
F
OUTBD
QTY
BRIGHTNESS
EXTERNAL
TEST
DECR

STATUS

AUX FUEL QTY
POUNDS

DECR

TOTAL

LEFT

NO
FLOW

INCR

9990

OUTBD

RIGHT

INCR

INBD

VENT
SENSOR

INBD

EMPTY

OUTBD

EMPTY

FAIL
DEGRADED

OVFL

CAL

AA0665
SA

Figure 4-27. Auxiliary Fuel Management Control Panel
2. TEST button - Release. Digits should display
8s in sequence from left to right three times; 5
seconds later, display GOOD or EO failure
code; 3 seconds later, display type fuel density,
then fuel TOTAL.

number, disregarding the numbers to the
right of the decimal point.
5. INCR/DECR switch - Set calibration.
6. Auxiliary fuel quantity switch - INBD.

3. Deleted.
4. Auxiliary fuel quantity switch - CAL.
NOTE
CAL is the calibration value marked on the
fuel flow transmitter. Enter the four digit

4-54

Change 10

7. INCR/DECR switch - Set inboard fuel quantity.
8. Auxiliary fuel quantity switch - OUTBD.
9. INCR/DECR switch - Set outboard fuel quantity.

TM 1-1520-237-10

10. Auxiliary fuel quantity switch move to TOTAL
- Check.
11. PRESS OUTBD and INBD switches - As desired.

TOTAL AUXILIARY FUEL REMAINING
(BASED
ON
JP-4 DENSITY)

TRANSFER
START WHEN
ONE
MAIN
FUEL
TANK
QUANTITY
LESS THAN

TRANSFER
STOP
WHEN
EACH
MAIN
FUEL
TANK
QUANTITY
MORE THAN

8840-7041 lbs

950 lbs

1000 lbs

7040-5001 lbs

750 lbs

1000 lbs

5000-0 lbs

600 lbs

1000 lbs

4.22.6 Fuel Transfer Sequence.

CAUTION

Fuel transfer sequence must be carefully
planned and executed in order to maintain CG within limits.
Fuel transfer sequence shall be based on mission requirement and center of gravity limitations. Automatic
transfer is started when the proper switches are manipulated
and fuel level is as shown below and external range tanks
internal pressure is increased enough to force fuel to the
main tanks. Transfer will continue until the main tank signal conditioner provides a signal through the microprocessor to stop fuel transfer. This cycle is done as required until
interrupted by placing the MODE switch to OFF or
MANUAL or placing the PRESS switch OFF. Manual
transfer will be started on selection of MANUAL and appropriate switches, and external fuel tanks are bleed-air
pressurized to start fuel transfer from external tank(s) to
main tanks. Transfer will continue until tanks are full. They
will remain full as long as the manual mode remains engaged. Manual transfer requires close monitoring of fuel
level to initiate and stop transfer to remain within CG limits. The automatic transfer sequence is as follows:

4.22.7 External Extended Range Fuel Transfer
Check.
NOTE
When ambient temperature is below 4°C
(40°F), ESSS/ERFS shall not be turned off
after transfer check has been completed to
avoid potential for freeze-up of the pressure
regulator.
1. AIR SOURCE HEAT/START switch - ENG.
2. FUEL BOOST PUMP CONTROL switches Check ON.

WARNING
FUEL BOOST PUMP CONTROL
switches shall remain on during external
range fuel transfer and remain on for 10
minutes after PRESS switches are moved
to OFF. Failure to observe this warning
may cause engine flameout.
3. PRESS OUTBD and INBD switches - ON for
tanks installed.
4. Fuel quantity switch - TOTAL.
5. TANKS switch - OUTBD.
6. MODE switch - MANUAL.

Change 10

4-55

TM 1-1520-237-10

7. MANUAL XFR RIGHT switch - ON.
8. Main FUEL QTY TOTAL FUEL readout Check for increase of about 20 pounds.

4.22.7.2 External Extended Range Fuel Transfer In
MANUAL Mode.
If AUTO mode is inoperative, transfer in MANUAL
mode as follows:

9. TANKS switch - Repeat steps 7 and 8 for
INBD, if installed.
CAUTION

10. MANUAL XFR RIGHT switch - OFF.
11. MANUAL XFR LEFT switch - ON.
12. Repeat steps 8. and 9. for MANUAL XFR
LEFT.

1. AIR SOURCE HEAT/START switch - ENG.

13. MANUAL XFR LEFT switch - OFF.

2. FUEL BOOST PUMP CONTROL switches Check ON.

14. External extended range fuel system - Set as
desired.
4.22.7.1 External Extended Range Fuel Transfer In
AUTO Mode.
NOTE
If either main fuel quantity is below 1,000
lbs., selecting the automatic mode may initiate a transfer sequence.
Allow sufficient time for tank pressurization
(approximately 10 minutes for a half full
230-gallon tank).
During transfer, periodically verify that
AUXILIARY FUEL MANAGEMENT
panel quantity is decreasing at a minimum
of 40 pounds per minute, per tank pair. Fuel
transfer rate of less than 40 pounds per
minute may indicate reduced flow from one
or both tanks.

4-56

Monitor fuel transfer to remain within
CG limits and avoid asymmetric loading.

3. PRESS OUTBD and INBD switches - ON for
tanks installed.
4. MODE switch - MANUAL.
5. TANKS switch - OUTBD. Switch to INBD
after outboard tanks are empty.
6. MANUAL XFR switches RIGHT and LEFT
- ON.
4.22.7.3 External Extended Range Fuel Flow Verification In Manual Mode. If extended range without
landing is required and the aircraft is not equipped with an
ERFS fuel indicating system, verify fuel flow from each
tank as follows:
NOTE
Ensure main fuel tanks are not completely
full.
1. AIR SOURCE HEAT/START switch - ENG.

1. AIR SOURCE HEAT/START switch - ENG.

2. FUEL BOOST PUMP CONTROL switches Check ON.

3. PRESS OUTBD and INBD switches - ON for
tanks installed.

4. MODE switch - MANUAL.

4. MODE switch - AUTO.

5. TANKS switch - OUTBD. Switch to INBD
after outboard tanks are empty.

6. MANUAL XFR RIGHT switch - ON. Note
the rate of decrease of the AUX FUEL QTY

Change 10

TM 1-1520-237-10

POUNDS indicator. The normal transfer fuel
flow rate per tank should be between 20 to 38
pounds per minute.
7. MANUAL XFR RIGHT switch - OFF.
8. Repeat steps 6. and 7. for left tank.

4.22A EXTERNAL AUXILIARY FUEL MANAGEMENT SYSTEM. AFMS
The external extended range fuel system is supported by
the external stores support system. The 230-gallon and 450gallon jettisonable tanks may be suspended from the vertical stores pylons (VSP). Removable fuel lines, bleed-air

Table 4-3. Extended Range Fuel System Degraded Operation Chart ERFS
SYSTEM FAILURE

DEGRADED

AUX FUEL

DESCRIPTION OF

CODES AND

LIGHT

CAUTION

DEGRADED OPERATION

INDICATIONS
E01 MICROPROCESSOR ERROR
E03 FLOWMETER DISCONNECTED
E04 ERROR FUEL FLOW CIRCUITS
E05 ERROR FUEL FLOW
COMPUTATION
E06 MEMORY ERROR

E02 TEMPERATURE SENSOR NOT
CONNECTED OR OUT OF RANGE

LOSS OF
DIGITAL READOUT

LOSS OF ONE MAIN TANK LEVEL
QUANTITY SENSOR OR LOSS OF
ONE SIGNAL CONDITIONER INPUT
FAILED AUX TANK EMPTY SENSOR
PROVIDES FALSE EMPTY SIGNAL

LIGHT

OFF

ON-IF FUEL
TRANSFER
SELECTED

1. AUTO XFR CAPABILITIES
REMAIN
2. DEFAULTS TO CURRENT XFR
SCHEDULE
3. PILOT MUST COMPUTE FUEL
USAGE
1. AUTO XFR CAPABILITIES
REMAIN
2. PERFORMING TWO SELF-TESTS
WILL:
A. CLEAR FAILURE CODE AND
REGAIN FUEL REMAINING INFO
B. RESET AUX FUEL LIGHT
C. DEFAULT TO PRESELECTED
TEMP VALUE
1. AUTO XFR CAPABILITIES
REMAIN
2. NO FLOW AND EMPTY
MONITORING INDICATIONS
REMAIN
3. PILOT MUST COMPUTE FUEL
USAGE
NO DEGRADATION
AUX TANK FUEL TRANSFER
SHUTOFF VALVE CLOSES.
PILOT SELECTING MANUAL
MODE REOPENS VALVE.

Change 4

4-57

TM 1-1520-237-10

lines, valves, and electrical connectors are within the horizontal stores supports (HSS). A tank pressurizing system,
using bleed-air, transfers fuel from the external tanks to the
main tanks. Fuel lines carrying fuel to the No. 1 and No. 2
main fuel tanks contain check valves to prevent backflow.
The external tanks are gravity refueled only. Control of the
system is provided by a auxiliary fuel management panel
(AFMP) located in the center of the instrument panel. Dimming control for the AFMP panel lighting is provided by
the cockpit INST LT NON FLT rheostat on the upper
console. Dimming control for all fuel quantity displays and
annunciators on the AFMP is provided by the LIGHTED
SWITCHES rheostat on the upper console only when the
caution/advisory panel is in the DIM mode. Power for the
auxiliary fuel management system is provided from the No.
2 dc primary bus through circuit breakers marked EXT
FUEL RH and NO. 2 XFER CONTROL on the mission
readiness circuit breaker panel.
4.22A.1 External Auxiliary Fuel Management System. This system contains two or four tanks suspended
from supports outboard of the fuselage. The tanks contain
baffles to prevent fuel sloshing. Quick-disconnect valves
are provided in external fuel and bleed-air lines to provide
seals when tanks are jettisoned or removed. If tanks are not
installed, the fuel quantity display for the removed tank is
blank when the XFER MODE is OFF. When the XFER
MODE switch is in any other position, the removed tank
display will show NT.
4.22A.2 Auxiliary Fuel Management Control Panel
(AFMP). The AFMP (Figure 4-27.1) contains all controls
for operating the external extended range fuel system. Controls description is as follows:

CONTROL/
INDICATOR

FUNCTION

PRESS switch

Provides control of bleed air
pressurization of auxiliary tanks.

CONTROL/
INDICATOR
OFF

Opens bleed-air valves to all
installed tanks for pressurization.

OUTBD

Opens bleed-air valves to outboard
tanks for pressurization.

INBD

Opens bleed-air valves to inboard
tanks for pressurization.

4-58

Change 4

Closes bleed-air valves to tanks.

XFER FROM
INBD

Selects fuel transfer from inboard
tanks.

OUTBD

Selects fuel transfer from outboard
tanks.

XFER MODE

Selects AUTO-OFF-MAN mode
of fuel transfer from external fuel
tanks.

AUTO

Activates automatic fuel transfer.

OFF

Closes all fuel transfer valves.
Interrupts automatic or manual
transfer mode of operation.
Deactivates NO FLOW indicators.

MAN

Selects manual transfer mode.
Activates MAN XFER switches.

MAN XFER
LEFT

Transfers from left inboard or
outboard tank.

BOTH

Transfers from both left and right
inboard or outboard tanks.

RIGHT

Transfers from right inboard or
outboard tank.

AUX FUEL QTY
LBS
L OUTBD

Pounds of fuel remaining in the left
outboard tank to the nearest 10 lbs.

L INBD

Pounds of fuel remaining in the left
inboard tank to the nearest 10 lbs.

R OUTBD

Pounds of fuel remaining in the
right outboard tank to the nearest
10 lbs.

R INBD

Pounds of fuel remaining in the
right inboard tank to the nearest 10
lbs.

PRESS
ALL

FUNCTION

TM 1-1520-237-10

TEST /
RESET

AUX FUEL
NO
FLOW

OUTBD
EMPTY

VENT
FAIL

QTY LBS
IMBAL

VENT
OVFL

INBD

EMPTY

NO
FLOW

OUTBD
R

EMPTY

XFER MODE MAN XFER XFER FROM PRESS
AUTO

LEFT

INBD

B
O
T
H

O
F
F
MAN

ALL
OUTBD
INBD

RIGHT

OUTBD

OFF

AB0820
SA

Figure 4-27.1. Auxiliary Fuel Management Control Panel

CONTROL/
INDICATOR

FUNCTION

TEST/RESET

On the ground (weight on wheels) Activates the Initiated Built in Test
(IBIT) and conducts a lamp test of
all displays and indicator lights.
System malfunctions are displayed
from left to right in the fuel quantity displays, using the error codes
(Table 4-4). Fuel quantity displays
return after the lamp test, if no errors are identified. Acknowledges
E04 through E07 error codes.

CONTROL/
INDICATOR

FUNCTION
The NO FLOW condition must exist for 5 seconds before the AUX
FUEL caution light and MASTER
CAUTION lights are activated.
One of these indications will occur
if a tank empty sensor fails (for related tank).
vA false EMPTY light (tank shows
quantity greater than zero).
vA false NT fuel quantity indication (tank installed, AUTO or
MAN mode selected).

In-flight (weight off wheels) - Resets AUX FUEL caution light and
MASTER CAUTION but does not
correct the condition.

vA false 9blank9 fuel quantity indication (tank installed, XFER
MODE is OFF).

NOTE
Illumination of annunciators on the
AFMP will activate AUX FUEL
caution light and MASTER CAUTION lights.

AFMS

vDegraded operation for the above
conditions: AUTO mode is disabled, use MAN mode.
AUX FUEL QTY
LBS

Change 4

4-59

TM 1-1520-237-10

CONTROL/
INDICATOR

FUNCTION

*R NO FLOW
light

No fuel flowing from right tank(s).

*R INBD EMPTY
light

Right inboard tank fuel exhausted.

*R OUTBD
EMPTY light

Right outboard tank fuel exhausted.

*L NO FLOW
light

No fuel flowing from left tank(s).

*L INBD EMPTY
light

Left inboard tank fuel exhausted.

*L OUTBD
EMPTY light

Left outboard tank fuel exhausted.
NOTE
EMPTY light will only activate
AUX FUEL caution light and
MASTER CAUTION lights when
its tank is selected for transfer.
If the EMPTY light is on for more
than 10 seconds and the fuel quantity display is greater than zero, the
tank empty sensor has failed.

VENT FAIL lights

Vent sensor inoperative.

*VENT OVFL
lights

Indicates when AFMS detects the
presence of fuel in vent line.

* IMBAL light

Indicates a lateral imbalance of 400
to 680 pounds between fuel quantity indications for outboard tanks.

4.22A.3 External Auxiliary Fuel
Quantity Indicating System. AFMS

Management

NOTE

The AUX FUEL QTY LB digital readout indicates
(Figure 4-27.1) the amount of fuel in each of the installed
external tanks in ten pound increments. A fuel probe in
each of the external tanks sends a signal to the AFMP proportional to the fuel level in the tank. The level of fuel at
the probe varies significantly with tank attitude. Aircraft
body pitch angle data from the copilot’s vertical gyro is
used by the AFMP so that the fuel quantity displayed includes compensation for angles between 10° nose down
and 1° nose up. E07 will result in a drop in AUX fuel
quantities for pitch down attitudes. A pitch up attitude may
result in a slight increase. The copilot’s attitude indicator
will not indicate OFF if the E07 error code occurs due to
failure of AFMP circuits external to the gyro. The pilot’s
vertical gyro is not connected to the AFMP and therefore
not available for attitude compensation. Level flight provides most accurate fuel quantity. A filter is incorporated in
the AFMP software to minimize fuel quantity variations
due to fuel slosh during maneuvering flight.
NOTE
The AFMP may display fuel quantities of up
to 150 pounds below the actual fuel quantity
per tank due to tank angle when the helicopter is on the ground.
A lateral imbalance is defined as any difference in fuel quantity readings between tank
pairs greater than 400 pounds. The crew
should not wait for the illumination of the
IMBAL light to begin correcting the lateral
imbalance condition.
The AFMP has three built in test (BIT) functions; power
up (PBIT), initiated (IBIT), and continuous (CBIT). PBIT
and IBIT are disabled in-flight by the WOW interlock. Circuits tested and related error codes are shown in Table 4-4.
PBIT is initiated when power is applied or reapplied to the
AFMP. IBIT is initiated by pressing the TEST/RESET
button for one second or more. CBIT is activated at the
completion of PBIT or IBIT and runs continuously.
NOTE

Unmodified 230-gallon and 450-gallon tanks
are prohibited from use on helicopters modified for AFMS. Crews should visually inspect each tank identification plate to verify
that only AFMS modified tanks are installed
on AFMS modified helicopters.

4-60

Change 10

The transition from APU to main generators
during engine run up can cause the PBIT to
be initiated. The AFMP will display fuel
quantities once the PBIT functions are completed.

TM 1-1520-237-10

4.22A.4 External Auxiliary Fuel Management System Modes of Operation.
4.22A.4.1 Automatic Mode Fuel Transfer.

WARNING
FUEL BOOST PUMP CONTROL
switches shall remain on during external
range fuel transfer and remain on for 10
minutes after PRESS switch is moved to
OFF. Failure to observe this warning may
cause engine flameout.

CAUTION

Monitor fuel transfer to remain within
CG limits and avoid asymmetric loading.
The fuel transfer sequence shall be based on mission
requirements and center of gravity limitations. Automatic
transfer is controlled by the AFMP sensing the No. 1 and 2
main tank fuel quantities to start and stop fuel transfer.
When AUTO mode is selected, transfer starts when the fuel
level in either main tank falls below 1,000 pounds, and the
external tanks are pressurized. The NO FLOW lights may
flicker upon initiation of fuel transfer. The AFMP transfers
fuel from both tanks selected with the XFER FROM
switch regardless of the position of the MAN XFER
switch. Fuel transfer stops when the TOTAL FUEL quantity indicates 2,200 pounds or when the EMPTY light on
either tank or VENT OVFL light illuminates on the AFMP.
4.22A.4.2 Manual Mode Fuel Transfer.

AFMS

CAUTION

Monitor fuel transfer to remain within
CG limits and avoid asymmetric loading.

If the IMBAL indicator illuminates, the
crew should verify selection of the heavy
tank and closely monitor the fuel quantity
displays on the AFMP. No additional
warnings are provided by the AFMP or
caution advisory system if the crew selects
the wrong tank with the MAN XFER
switch.
Manual transfer requires close monitoring of the main
fuel quantity and AFMP fuel quantity displays to remain
within CG limits and maintain lateral balance. Manual
mode is initiated by the pilot when the XFER MODE
switch is placed to MAN and external tanks are pressurized. Fuel transfer will continue as long as MAN is selected. The NO FLOW lights will randomly flicker as fuel
is transferred until the main fuel quantity reaches approximately 2,300 pounds, unless the XFER MODE switch is
placed to OFF or AUTO. The NO FLOW condition results when the tank pressurization can no longer add fuel to
the tanks due to activation of the high level fuel shutoff
valves in the main tanks. Illumination of the EMPTY light
alerts the pilot to change tank pairs using the XFER FROM
switch, select another tank using the MAN XFER switch,
or place the XFER MODE switch to OFF. To avoid
pumping air into the main tanks, do not wait for the NO
FLOW light to illuminate. Sloshing of the fuel will cause
frequent illumination of the EMPTY light only when the
tank is selected for fuel transfer. Sloshing of fuel and activation of the fuel transfer valves using the XFER MODE
switch can cause flickering of the NO FLOW light; however the AUX FUEL and MASTER CAUTION will only
illuminate if the NO FLOW lights remain illuminated for
more than 5 seconds. If lateral imbalance results during fuel
transfer, the AFMP senses the imbalance and illuminates
the IMBAL light. The lateral imbalance can be resolved by
selecting fuel transfer from the heavy tank. The IMBAL
light illuminates with approximately 685 pound difference
between outboard tanks, and will remain illuminated until
the lateral imbalance is reduced below approximately 450
pound difference between outboard tanks.
4.22A.5 Auxiliary Fuel Management Control Panel
- TEST. AFMS
1. AUX FUEL QTY digital readouts - Note current reading.
2. TEST/RESET button - Press. All annunciators
will light and E07 will be displayed, if test is
initiated within approximately one minute of
applying ac power. If E07 appears, press
TEST/RESET again. E07 should not appear
until test is finished.

Change 10

4-60.1

TM 1-1520-237-10

3. BIT indications - Check. Fuel quantity displays
will appear upon completion of BIT. The pitch
attitude sensor error E07 test has a 5 minute
delay to allow the vertical gyro to obtain full
operating speed.

10. XFER MODE switch - OFF.
11. External extended range fuel system - Set as
desired.
4.22A.6.1 Fuel Transfer in AUTO Mode. AFMS

4. PRESS switch - As desired for tanks installed.
4.22A.6 External Auxiliary Fuel Management System Fuel Transfer Check. AFMS
NOTE
When ambient temperature is below 4° C
(39° F), ERFS/AFMS shall not be turned off
after transfer check has been completed to
avoid potential freeze up of the pressure
regulator.
1. AIR SOURCE HEAT/START switch - ENG.
2. FUEL BOOST PUMP CONTROL switches Check ON.

NOTE
Allow sufficient time for tank pressurization
(approximately 10 minutes for a half full 230
gallon tank).
During transfer, periodically verify the TOTAL FUEL quantity remains above 2,000
pounds and the selected tank pair remains in
balance. A decrease below 2,000 pounds on
the TOTAL FUEL quantity display or the
generation of an imbalance in the AUTO
mode may indicate reduced flow from one
or both of the external tanks selected.
1. AIR SOURCE HEAT/START switch - ENG.

2. FUEL BOOST PUMP CONTROL switches Check ON.
3. PRESS switch - As required for tanks installed.
4. XFER FROM switch - OUTBD, then INBD.
5. XFER MODE switch - AUTO.
4.22A.6.2 Fuel Transfer in MANUAL Mode. AFMS
If the AUTO mode is inoperative or a lateral imbalance
greater than 400 pounds between tank pairs is identified,
transfer in the MAN mode as follows:

5. MAN XFER switch - BOTH.
CAUTION

6. XFER MODE switch - MAN.
7. Main FUEL QTY TOTAL FUEL readout Check for increase of approximately 40 pounds
and AUX FUEL QTY LBS decrase 20 pounds
per tank.
8. XFER FROM switch - INBD if installed.
9. Main FUEL QTY TOTAL FUEL readout Check for increase of approximately 40 pounds
and AUX FUEL QTY LBS decrase 20 pounds
per tank.
4-60.2

Change 10

Monitor fuel transfer to remain within
CG limits and avoid asymmetric loading.
1. AIR SOURCE HEAT/START switch - ENG.
2. FUEL BOOST PUMP CONTROL switches Check ON.
3. PRESS switch - As required for tanks installed.
4. XFER FROM switch - OUTBD, then INBD.

TM 1-1520-237-10

5. MAN XFER switch - BOTH or select heavy
tank to correct an imbalance.
6. XFER MODE switch - MAN.
4.23 EXTERNAL
(ESSS). ES

STORES

SUPPORT

SYSTEM

ESSS provides a means of carrying a variety of external
stores, including external extended range fuel tanks. The
ESSS consists of fixed and removable provisions.

lage strut support fittings for attaching two struts for each
HSS. In addition to exterior components, fixed provisions
are: interior helicopter provisions, including electrical harnesses, fuel lines, bleed-air lines, and circuit breakers.
4.23.2 External Stores Removable Provisions. The
external stores removable subsystem extends horizontally
from each side of the helicopter at station 301.5, buttline
42.0. Extending below each horizontal stores support (HSS)
are two vertical stores pylons (VSP) and attaching ejector
racks. The racks are used to attach fuel tanks or other external stores dispensers.

4.23.1 External Stores Fixed Provisions. Fixed provisions are: upper fuselage fixed fittings for attaching the
horizontal stores support (HSS) subsystem, and lower fuseTable 4-4. Auxiliary Fuel Management System Fault Messages AFMS
SYSTEM FAILURE CODES AND/
OR INDICATIONS

DESCRIPTION OF
DEGRADED OPERATIONS

E01 - AFMP microprocessor fail
E02 - AFMP memory fail
E03 - AFMP display fail

1. E01-E03 error codes displayed continuously if failure
occurs during PBIT or IBIT on the ground (WOW only).
2. Only MAN mode for transfer is available.

E04 - AFMP tank gaging electronics failure
E05 - Auxiliary tank probe (OPEN)
E06 - Auxiliary tank probe (SHORT)

1. Error codes display during PBIT or IBIT if failure
occurs on the ground (WOW only).
2. Acknowledge failure by pressing TEST/RESET
button and error code changes to the CBIT display (---)
or FP.
3. Manual mode only is available for the tank pair
selected. Fuel quantity must be calculated for the tank
with the failure.

(--- ) AFMP tank gaging electronics failure.
FP - Auxiliary tank probe circuit failure (open or short).
NF - No tank detected.

1. (--- ) is the failure indication for error code E04 and
FP is the failure indication for E05 and E06 during
CBIT on the ground or in flight or during PBIT or IBIT
on the ground after the crew acknowledges the error
using TEST/RESET.
2. Manual mode only is available for the tank pair with
the failure. Fuel quantity must be calculated for the tank
with the failure.
3. No tank detected by system, when that tank is
selected for transfer.

E07 - AFMP memory fail

1. May occur on the ground if IBIT is initiated within
approximately 1 minute of applying ac power to the
aircraft. Acknowledge E07 with the TEST/RESET
button.
2. If E07 is observed in flight or persists on the ground,
AUTO and MAN mode are available but the pitch
attitude correction defaults to a level attitude regardless
of flight condition.

Change 10

4-60.3

TM 1-1520-237-10

4.23.3 ESSS Side Position Lights. A position light is
on each outboard end of HSS. Those lights use the power
source provided to operate the standard installed position
lights, colors are the same. Upon installation of the HSS,
the electrical connectors connected to the jumper plugs,
providing power for the standard position lights, are removed and reconnected to the connectors from the HSS
position lights. Operation and power source for the ESSS
position lights are the same as for the standard installed
position lights.
4.23.4 External Stores Jettison Control Panel. The
jettison control panel (Figure 4-28) provides the capability
of phase jettison of all external stores or symmetrical jettison of fuel tanks. Interlock circuitry prevents jettison of
fuel tanks other than in pairs. Emergency jettison is completely independent of the primary jettison subsystem.

WARNING
The BRU-22A/A and MAU-40/A Ejector
Rack CARTRIDGES are explosive devices and must not be exposed to heat,
stray voltage or static electricity. Refer to
TM 9-1300-206 for information concerning handling and storage of ammunition.
The jettison control panel (Figure 4-28) contains all controls for jettisoning external stores. Jettison controls are as
follows:

CONTROL

OFF

Applies 28 volts from essential dc
bus to all stores stations when the
helicopter weight is off the wheels,
regardless of the rotary selector
switch. A 1-second time delay permits the outboard stations to jettison before the inboard stations.

Rotary selector
switch

Determines which station receives
primary jettison signal.

Change 10

FUNCTION
Prevents jettison signal from going
to any stores station.
NOTE

*If fuel tanks are connected to the
left, right, or both stores stations,
the BOTH mode of jettison is automatically selected even if the selector switch is at L or R.
INBD
L

*Directs jettison signal to inboard
left station.

*Directs jettison signal to inboard
right station.

BOTH

Directs jettison signal to inboard
left and right stores stations.

OUTBD
L

*Directs jettison signal to outboard
left station.

*Direct jettison signal to outboard
right station.

BOTH

Directs jettison signal to outboard
left and right stores stations.

ALL

Directs primary jettison signal to all
stores stations. Outboard stores will
jettison and 1 second later inboard
stores will jettison.

JETT

Applies 28 volts from primary dc
bus through the rotary selector
switch to the selected stores station
if the weight is off the wheels and
the selector switch is not OFF.

FUNCTION

EMER JETT ALL

4-60.4

CONTROL

TM 1-1520-237-10

4.24 RAPPELING ROPE CONNECTORS.
EMER
JETT
ALL

Rappeling rope connectors consist of four cabin ceiling
tie down fittings.

STORES JETTISON
JETT
INBD
BOTH

OUTBD
BOTH

R
L
OFF

4.25 MEDICAL EVACUATION (MEDEVAC) KIT.

R
L

WARNING

ALL

AA0664
SA

Figure 4-28. Stores Jettison Control Panel

4.23.5 Stores Jettison Control Operation.

CAUTION

To prevent unintentional jettison of external stores when the helicopter weight is
on the wheels, do not actuate any jettison
switch.
The jettison system provides two modes of jettisoning
external stores, primary and emergency. The primary subsystem uses the rotary selector switch and the JETT toggle
switch. The emergency jettison subsystem uses only the
EMER JETT ALL toggle switch. Primary jettison is used
when selective jettison is desired. The rotary switch is used
to select the stores point for release, and the JETT toggle
switch is used to actuate the release. Emergency jettison is
used to release all external stores through one actuation of
the EMER JETT ALL toggle switch, regardless of rotary
switch position. During primary (rotary switch ALL selected) and emergency jettison, a 1-second delay is provided after the outboard stores are released, before the inboard stores will be released. When one pair of tanks is
jettisoned in a four tank system, cccc will appear on the
AUX FUEL QTY POUNDS digital readout when the corresponding fuel quantity position is selected. The fuel remaining in the tanks jettisoned will be subtracted from the
total displayed when TOTAL is selected. Power to operate
the primary jettison subsystem is from the No. 1 dc primary
bus through circuit breakers marked ESSS JTSN INBD
and OUTBD. The emergency jettison subsystem is powered from the dc essential bus through circuit breakers
marked ESSS JTSN INBD and OUTBD.

Use of the medevac pedestal ambulatory
configuration for transport of personnel
other than patients or essential medical
personnel is prohibited.
A medevac kit consisting of a pedestal support assembly
and provisions for three rear-facing troop seats may be installed in the UH-60 helicopter (Figure 4-29) after removing the existing troop seats. The medevac pedestal assembly, when installed, is directly below the main transmission.
The pedestal can be turned about a vertical axis. Litter supports are cantilevered from the pedestal. The litter supports
may be positioned to accept four to six litter patients, up to
six ambulatory patients or essential medical personnel, or
combination thereof. The pedestal should be positioned
along the longitudinal axis of the helicopter for flight, to
provide maximum crash attenuation. The pedestal contains
restraint belts for each litter, restraint lap belts for each
ambulatory occupant, eight individually operated lights for
the four-man litter configuration, provisions for eight 1000
ml. intravenous fluid bags, and provisions for two size D
oxygen bottles. Another feature of the medevac kit is a 115
vac, 60 Hz frequency converter to provide electrical power
for use of standard hospital equipment. On missions not
requiring electrical power, the power pack may be left out.
The three-man rear-facing seat provisions are in the forward portion of the cabin, and accommodate standard troop
seats. The four man litter configuration allows rotation of
the pedestal so that the litter patients can be loaded from
either side of the helicopter. The six-man litter configuration also allows for side loading: however, the pedestal
must be rotated back to the locked position along the longitudinal axis of the helicopter after four litters are loaded.
Floor restraints are then installed to the cabin floor tiedown
studs on both sides of the pedestal. The last two litters are
placed on both sides of the pedestal between the floor restraints and secured. Only the upper supports are capable of
being tilted for loading or unloading of the litters. Unloading the patients is the reverse of loading. To convert to the
six-man ambulatory patient or essential medical personnel
configuration, the upper litter supports are folded down to
accommodate three patients or essential medical personnel
seated side by side on either side of the pedestal. The mede-

Change 10

4-60.5

TM 1-1520-237-10

vac pedestal ambulatory configuration provides significantly less crashworthiness capability (energy attenuation
and occupant restraint) than the troop seats.

ward pedestal, until pivot shaft is fully inserted into pivot
shaft hole on pedestal and handle lock is engaged.
(4) Repeat step (3) for other end of litter support.

4.25.1 Litter Support. Each litter support is attached to
the center pedestal by two end pivot shafts, (Figure 4-29)
and by two T-shaped fittings, which allows removal, interchange, or repositioning of the supports. Crashload absorption works on the deformation principal. There are five
pivot shaft support holes on the right and left side of the
center console at both ends. Behind the holes are support
rollers for the pivot shafts. From top to bottom, the top hole
is provisions for the upper litter in the six-litter configuration. The second hole is for the upper litter support of a
four-litter configuration. These end holes line up with a
central pivot shaft on the litter support. Only this litter position allows midposition pivoting for loading or unloading.
The third hole is for the center litter of the six-litter configuration. The fourth hole is used when installing the litter
support in the four-litter configuration. The third, fourth,
and fifth positions do not provide a tilt function.
4.25.2 Litter Lighting. Two litter lights are installed in
the pedestal at each litter (Figure 4-29). Each light contains
a PUSH-ON, PUSH-OFF switch. The positioning of those
lights is adjustable. Power to operate the litter lights is from
the No. 1 and No. 2 dc primary buses through circuit breakers on the mission readiness circuit breaker panel, marked
NO. 1 LTR LTS and NO. 2 LTR LTS. The lights are
operated from a split bus to provide one light at each litter
in case of a single dc primary bus failure.
4.25.3 Litter Support Installation. The upper litter
supports are supported by a center pivot shaft and two end
pivot shafts, one at each end of the support (Figure 4-29).
To tilt the upper end of the support only for loading or
unloading of litter patients, the center shaft remains locked
to the pedestal and the end shafts are disengaged for support pivoting. This system was designed to pivot about the
center shaft. Although the supports may be pivoted at either
end, more effort is required when the loaded litter is installed. To install the litter supports, do this:

b. Upper Litter Support Installation.
(1) Prepare support. Before installation, each center
pivot pin must be unlocked and retracted, and the handle
disengaged from its retainer. End pivot handles must be in
disengaged position.
(2) Tilt outer edge of litter support slightly down and
engage T-bars into split retention fittings at second support
hole from top of pedestal.
(3) Raise outer edge of litter support until support is
level.
(4) Insert end pivot shaft into pedestal by pulling on
pivot shaft lever lock, and moving lever toward pedestal
until end pivot shaft engages partway in end pivot support
hole.
(5) Position center pivot shaft lock handle counterclockwise to horizontal.
(6) Push center pivot shaft toward pedestal until shaft
is fully inserted into center pivot shaft hole. Opposite end
of litter support should be raised or lowered to help line up
center shaft on support with center hole on pedestal.
(7) Turn center pivot lock lever clockwise to horizontal.
(8) Repeat step (4) for other end of litter support. Now
slide both end pivot shafts in fully by moving pivot lever
lock handle to engage position.
c. Litter Support Installation for Ambulatory Patient
Seating.

a. Lower Litter Support Installation.
(1) Prepare support as in b(1) above.
(1) Before installation, each center pivot shaft must be
retracted and unlocked. The center pivot shaft handle must
be secured in the handle retainer. End pivot handles must
be in disengaged position.

(2) Engage T-bar on litter pan with split retention
brackets below support tilt stop brackets.

(2) Engage T-bars on litter support with split retention
fittings at bottom of pedestal.

(3) Position litter support at second from bottom litter
support end pivot hole on pedestal.

(3) Line up end pivot shafts with holes. Disengage
pivot shaft lever locks and move end pivot shaft lever to-

(4) Line up end pivot shafts with holes. Disengage
pivot shaft lever lock and move pivot shaft lever toward

4-60.6

Change 10

TM 1-1520-237-10

UPPER TROOP SEATS
SUPPORT

60 Hz FREQUENCY
CONVERTER

LITTER LIGHT
(TYPICAL 8)

CEILING
SUPPORT

LITTER RESTRAINT
BELTS (TYPICAL 8)

1
2
LITTER

TROOP
SEATS

TROOP SEATS
FLOOR PLATE

LITTER SUPPORT
(TYPICAL 4)

FLOOR
SUPPORT
PLATE

LITTER RESTRAINT
BELT FITTING (TYPICAL 8)

END PIVOT SHAFT
LEVER (TYPICAL 8)

AA0371_1
SA

Figure 4-29. Medevac and Seat System (Sheet 1 of 5)
pedestal, until pivot shaft is fully inserted into pivot shaft
hole on pedestal and handle lock is engaged.
(5) Repeat step (3) for other end of litter support.
4.25.4 Litter Support Removal. Removal of the litter
support is the reverse of installation. Before removal, any
litters on the support should be removed and belts unlocked.
If IV or oxygen is installed, make certain hoses are not
tangled with supports, then proceed as required.
4.25.5 Medevac Seats Installation. The seat installation consists of three of the troop seats that were removed
for medevac system installation. Install required number of
seats at station 271.0.

4.25.6 Litter Loading and Unloading. Litters can be
loaded and unloaded laterally, directly onto the litter supports, from either side of the helicopter. Whenever rescue
hoist and medevac kit are installed simultaneously, the upper, right litter support should be removed from the aircraft.
The lower, right support may be stowed if not actually in
use. The lower right litter support shall be installed in the
lowest position and used when transporting more than two
litter patients or when conducting hoist operations with a
stokes litter. Loading of a stokes litter patient may be facilitated by rotating the litter pedestal approximately 30 degrees from the fly position. When returning the pedestal to
the fly position the aft right corner of the litter support must
be lifted to prevent interference with the lower hoist mount

4-61

TM 1-1520-237-10

LOWER LITTER
IV BAGS

PRESSURE
GAGE

UPPER LITTER
IV BAGS

OXYGEN
TANK

FLOW
GAGE
OXYGEN
REGULATOR

OXYGEN TANK
SHUTOFF VALVE

3
4
5
6
7
8
9
H4
IC636

OXYGEN
HUMIDIFIER

LITTER
TIEDOWN
STRAP

ROTATION
RELEASE
LOCK HANDLE
OXYGEN TANK
RESTRAINT STRAP
CENTER
PEDESTAL

TYPICAL IV/OXYGEN TANK INSTALLATION
(SAME AS OTHER END OF PEDESTAL)

1
A

OXYGEN TANK

OXYGEN REGULATOR

OXYGEN TANK
SHUTOFF VALVE

OXYGEN TANK
RESTRAINT STRAP

FLOW GAGE
PRESSURE GAGE

ALTERNATE OXYGEN REGULATOR INSTALLATION

Figure 4-29. Medevac and Seat System (Sheet 2 of 5)
4-62

AA0371_2
SA

TM 1-1520-237-10

UPPER SUPPORT PIVOT HOLE
(PROVISIONAL 6 LITTER)

LITTER FEET
WOOD STOPS

SUPPORT RESTRAINT
SPLIT GUIDE

UPPER SUPPORT END
SHAFT HOLE (4 LITTER)
LITTER SUPPORT
TILT STOP BRACKET
GUIDE PLATE

CENTER SUPPORT
END SHAFT HOLE
(PROVISIONAL)

LITTER SUPPORT
HOLE AMBULATORY
PATIENT SEAT
(EMERGENCY)
LOWER SUPPORT END SHAFT
HOLE (4 LITTER)

LOWER LITTER SUPPORT
RESTRAINT BELT

LITTER PIN IN LOAD UNLOAD (TILT) POSITION
(SAME AT OTHER SIDE OF PEDESTAL)

2
AA0371_3
SA

Figure 4-29. Medevac and Seat System (Sheet 3 of 5)
bracket. To load and unload litter patients, assuming the
medevac kit is in the flight position (litters along longitudinal axis), do this:
1. Both cabin doors - Open.
2. Pedestal rotation lock release handle - Pull
handle and turn pedestal clockwise 90° (viewed
from above). On helicopters with extended external range fuel tanks installed, the pedestal
will rotate only 60° from center line for loading
litter patients.

3. Release lock handle while turning pedestal.
Pedestal will automatically lock in a lateral position for loading and unloading.
4. Release both litter support end pivot shaft on
upper litters. Disengage pivot lever locks and
move levers away from pedestal. Hold support
with opposite hand. Release lever. End pivot
shafts should rest on fitting at hole. Litter support is now ready to be loaded from either side.
Select side desired. Move end pivot release lever about 1 inch more to compress the shaft

4-63

TM 1-1520-237-10

SUPPORT STOWAGE
STRAP IN USE
LITTER SUPPORT
STOWAGE STRAP

STOWED LITTER
SUPPORT

UPPER STOWAGE
ASSEMBLY

SUPPORT STOWAGE
STRAPS STOWED

LOWER STOWAGE
ASSEMBLY

STOWAGE
ASSEMBLY
PIN

LITTER SUPPORT STOWAGE

3
AA0371_4
SA

Figure 4-29. Medevac and Seat System (Sheet 4 of 5)

4-64

TM 1-1520-237-10

LEFT SIDE SHOWN

RIGHT SIDE SHOWN

STOWAGE ASSEMBLY STOWED

LITTER SUPPORTS STOWED
WITH CROSS STRAPS

4
AB0860
SA

Figure 4-29. Medevac and Seat System (Sheet 5 of 5)
Change 3

4-65

TM 1-1520-237-10

springs, which allows the shaft to clear the end
guide and the litter support to be lowered at the
end. During the lowering, release pivot shaft
lever to allow pivot shaft spring to push shaft
onto lower stop fitting.
5. Using two persons (one each side or end) -Place
litter with patient on end of upper support and
push litter into position. Note that litter feet
must be trapped between wood stops on litter
support. If three or more patients are to be
loaded, the upper supports must be loaded first.
The reverse applies to unloading.
6. To tilt upper litter support end, pull shaft lever
lock and move lever away from pedestal at support end which is being raised. Pivot litter support to level position until pivot shaft holes are
lined up with pivot shafts. Move levers toward
pedestal until shaft is fully inserted into shaft
holes and handle locks are engaged.
7. Lower litters - Using two persons (one each
side or end) place litter with patient on end of
support and push litter into position. Note that
litter feet must be trapped between wood stops
on litter support.
8. Litter straps - Extend straps (on pedestal) and
engage in buckle on litter supports. Pull straps
out uniformly to engage; partial pulling will require complete retraction of the belt to disengage belt lock.
9. Pedestal rotation lock handle - Pull and turn
pedestal counterclockwise 90° (viewed from
above) into flight position (longitudinal axis),
and release handle.
10. Cabin doors - Close.
11. Unloading is reverse of loading after litter
straps are removed and oxygen and IV tubes
are checked to make certain tangling will not
occur with litter or support.
4.25.7 IV Bags and Oxygen Tanks Installation.

General’s directives, and must have oxygen regulators attached.
Provisions for IV bags and oxygen tanks are on the top
of the pedestal at each end. Four IV bags may be attached
to each IV/oxygen assembly (Figure 4-29). IV bag hooks at
the outer end of the assembly are used for the lower litters
and the inner hooks are used for the upper litters. Eyelets at
the top of the bag are placed on the IV hooks and the bags
are hung downward. To prevent damage to IV bags, check
clearance between transmission drip pan drain tube clamps
and installed IV bags. Flow adjustment and replacement
will be done by the medical attendant. Oxygen tanks are
inserted into the assembly, bottom first. A restraint strap is
provided to prevent the tank from falling out during normal
maneuvering during flight. The strap is placed across the
regulator in a manner and routed as shown in Figure 4-29,
to prevent the restraint strap from slipping. The strap ends
are attached and drawn tight to keep the tank secure.
4.25.8 Litter Support Stowage.

WARNING
Storage of the litter support in the upper
level stowed position can be dangerous
during a crash sequence due to the release
of the litter support from the carousel.
Advise storage in this manner be avoided.
Maintain this litter support in the installed position or place in the back of the
carousel in the ambulatory level if there
are no occupants along the aft bulkhead
(Row 5).
The litter supports may be stowed along the center pedestal on each side, one above the other (Figure 4-29). Stowage brackets at each end of the pedestal provide lower support of the supports, and prevent the supports from moving
away from the pedestal. Web straps attached to rings are
used to hold the upper ends of the supports to the pedestal.
Pins are used to hold the stowage brackets in a stowed
position against the pedestal end. Two brackets are provided for each litter support. The top support must be
stowed first, then the lower support. For reinstallation the
sequence is reversed.

CAUTION

The pilot must be advised when oxygen is
on board, its use must be per the Surgeon

4-66

Change 7

1. Lower the stowage support arm to the horizontal position and insert the support arm stowage
pin through the support arm and into the center
pedestal.

TM 1-1520-237-10

NOTE
Improper positioning of the support arm
stowage pin reduces the holding capability
of the support arm, which may cause the
support arm to shear its pivot bolt during a
hard landing or aircraft mishap.
2. Place the litter pan in the stowed position, with
the top of the litter pan against the center pedestal and the pivot support arm properly
stowed.
3. Secure the litter pan to the center pedestal by
routing the opposite side web strap around the
upper portion of the litter pan handle. Secure
the metal clasp to the metal ring and tighten the
web strap. (Use of the opposite side web strap
will reduce excess movement of the litter pan
while stowed).
4. If only one upper litter pan is to be stowed, as
in step 3, additional security may be added by
routing the same side web strap around the
lower portion of the litter pan handle and fastening the web strap.

CAUTION

Do not store equipment between the
stowed litter pan and the center pedestal.
5. The lower litter pan will be stowed in the same
manner as in steps 1 through 3. The same side
web strap may be used to secure the lower portion of the litter pan as in step 4 if only one
lower litter is to be stowed.
6. Removal of stowed litter pans is accomplished
in the reverse order of steps 1 through 4.
4.26 APU INLET PARTICLE SEPARATOR (IPS) KIT
(HELICOPTERS WITH IPS KIT INSTALLED).
The APU IPS Kit provides APU inlet air filtration via a
centrifugal particle separator unit. The separator is attached
to the APU radial inlet housing and provides for collection
and overboard exhausting of scavenge particles. The passive separator operation employs APU bleed air to drive an
ejector pump used for particle scavenging. The IPS kit is

designed to be physically compatible with both HIRSS and
non-HIRSS helicopters with the T-62T-40-1 series 200/300
APU installations only. The kit consists of three categories
of removable components:
a. Air particle separator assembly.
b. APU modification kit - Parts required to modify the
APU to accept the separator assembly.
c. Airframe provisions - Parts required to install the
separator assembly and provide bleed air supply and scavenge exhaust provisions.
4.27 SNOW SKIS.
Landing gear skis are constructed of fiberglassreinforced plastic, and attached to the landing gear axle.
The skis have spring cylinders and check cables to retain
the ski in a 5° nose up attitude during flight.

CAUTION

Installation of skis requires removal of
landing gear wire cutters and severely degrades the aircraft wire strike capability.
Upon removal of skis, wire strike hardware shall be reinstalled, restoring aircraft to standard configuration prior to
next flight.
Cockpit entry/exit paths are partially restricted by the main skis making cockpit
entry/exit slightly more difficult. Additionally, the cabin entry/exit doors are
partially restricted making the loading/
unloading of cargo slightly more difficult.
NOTE
The hinged main gear ski shall only be used
on the right landing gear of helicopters
equipped with rescue hoists. The hinged ski
is equipped with a retraction cable. This
cable may be removed if it interferes with
the hoist, or other equipment, and alternative
retraction methods, such as a gaff, may be
used.

Change 9

4-67/(4-68 Blank)

TM 1-1520-237-10

CHAPTER 5
OPERATING LIMITS AND RESTRICTIONS
Section I GENERAL
5.1 PURPOSE.

restrictions for UH-60L helicopters 9626723 and subsequent.

This chapter identifies or refers to all important operating limits and restrictions that shall be observed during
ground and flight operations.
5.2 GENERAL.
The operating limitations set forth in this chapter are the
direct results of design analysis, tests, and operating experiences. Compliance with these limits will allow the pilot to
safely perform the assigned missions and to derive maximum use from the aircraft.

5.3 EXCEEDING OPERATIONAL LIMITS.
Any time an operational limit is exceeded an appropriate
entry shall be made on DA Form 2408-13-1. Entry shall
state what limit or limits were exceeded, range, time beyond limits, and any additional data that would aid maintenance personnel in the maintenance action that may be required. The helicopter shall not be flown until corrective
action is taken.
5.4 MINIMUM CREW REQUIREMENTS.

NOTE
See current Interim Statement of Airworthiness Qualification for operating limits and
restrictions for EH-60A helicopters.
EH

The minimum crew required to fly the helicopter is two
pilots. Additional crewmembers as required will be added
at the discretion of the commander, in accordance with pertinent Department of the Army regulations.

See current Interim Statement of Airworthiness Qualification for operating limits and

Change 6

5-1

TM 1-1520-237-10

Section II SYSTEM LIMITS
5.5 INSTRUMENT MARKING COLOR CODES.

up to 120% RPM R are authorized for use by maintenance
test flight pilots during autorotational RPM checks.

NOTE
5.7 MAIN TRANSMISSION MODULE LIMITATIONS.
Instrument/color markings may differ from
actual limits.
Operating limitations are shown as side arrows or colored strips on the instrument face plate of engine, flight and
utility system instruments (Figures 5-1, 5-2, and 5-3). Those
readings are shown by ascending and descending columns
of multicolor lights (red, yellow and green) measured
against vertical scales. RED markings indicate the limit
above or below which continued operation is likely to cause
damage or shorten component life. GREEN markings indicate the safe or normal range of operation. YELLOW
markings indicate the range when special attention should
be given to the operation covered by the instrument.
5.6 ROTOR LIMITATIONS.
It is not abnormal to observe a % RPM 1 and 2 speed
split during autorotational descent when the engines are
fully decoupled from the main rotor. A speed increase of
one engine from 100% reference to 103% maximum can be
expected. During power recovery, it is normal for the engine operating above 100% RPM to lead the other engine.
Refer to Figure 5-1 for limitations.
5.6.1 Rotor Start and Stop Limits. Maximum wind
velocity for rotor start or stop is 45 knots from any direction.
5.6.2 Rotor Speed Limitations. Refer to Figure 5-1
for rotor limitations. Power off (autorotation) rotor speeds

5-2

Change 3

a. Oil pressure should remain steady during steady state
forward flight or in level hover. Momentary fluctuations in
oil pressure may occur during transient maneuvers (i.e.
hovering in gusty wind conditions), or when flying with
pitch attitudes above +6°. These types of oil pressure fluctuations are acceptable, even when oil pressure drops into
the yellow range (below 30 psi). Oil pressure should remain
steady and should be in the 45 to 55 psi range for the
UH-60A/EH-60A, and 45 to 60 psi range for the UH-60L,
to ensure that when fluctuations occur they remain in the
acceptable range as defined above. If oil pressure is not
steady during steady state forward flight or in a level hover,
or if oil pressure is steady but under 45 psi, make an entry
on Form 2408-13-1. Sudden pressure drop (more than 10
PSI) without fluctuation requires an entry on Form 240813-1.
b. A demand for maximum power from engines with
different engine torque factors (ETF) will cause a torque
split when the low ETF engine reaches TGT limiting. This
torque split is normal. Under these circumstances, the high
power engine may exceed the dual engine limit. (Example:
#1 TRQ = 96% at TGT limiting, #2 TRQ is allowed to go
up to 104%. Total aircraft torque = (96%+104%)/2 =
100%).
c. With transmission oil temperature operation in the
precautionary range, an entry should be made on DA Form
2408-13-1 except when hovering in adverse conditions described in Chapter 8 Desert and Hot Weather Operations.

TM 1-1520-237-10

MAIN ROTOR
OVERSPEED
*

127%

137%

***

142%

% RPM

MAIN ROTOR % RPM R

TEST RTR OVERSPEED
*
***
**

POWER ON

130

1 R 2

TRANSIENT

101% − 107%

CONTINUOUS

95% − 101%

TRANSIENT

91% − 95%

130

120

110

ENGINE % RPM 1−2
105

105

100

101% − 105%

CONTINUOUS

95% − 101%

TRANSIENT

91% − 95%

12−SECOND
TRANSIENT

105% − 107%

TRANSIENT

POWER OFF (AUTOROTATION)

MAXIMUM

110%

TRANSIENT

105% − 110%

NORMAL

90% − 105%

0
1 R 2

AVOID OPERATIONS IN 20% − 40%
AND 60% − 90% RANGE EXCEPT
DURING START AND SHUTDOWN

FUEL
QTY
LB X 100

14
12
10

FUEL QUANTITY
NORMAL
PRECAUTIONARY

UH60A

200 − 1500 LBS

0 − 200 LBS

4
2
0
1

TOTAL
FUEL

FUEL
QTY
LB X 100

FUEL QUANTITY

UH−60L

12
10

NORMAL

200 − 1500 LBS

PRECAUTIONARY

0 − 200 LBS

8
6
4

LEGEND

2
0

RED

YELLOW
GREEN
DIGITAL READOUT

MAIN
FUEL

AA8670_1B
SA

Figure 5-1. Instrument Markings (Sheet 1 of 2)

Change 10

5-3

TM 1-1520-237-10

250
50

STAB
POS

S
T D
A E
B G

200

KNOTS

O
F
F

100

DEG

KIAS
LIMIT

30
40
DN

150

0o
10o
20o
30o
40o

150
100
80
60
45

AIRSPEED
MAXIMUM

193 KNOTS

REFER TO SECTION V FOR
ADDITIONAL AIRSPEED
LIMITATIONS

AA8670_2A
SA

Figure 5-1. Instrument Markings (Sheet 2 of 2)

5-4

TM 1-1520-237-10

Ng
SPEED
% X 10

11
10

ENGINE Ng

10−SECOND
TRANSIENT

102% − 105%

30−MINUTE
LIMIT

99% − 102%

CONTINUOUS

0 − 99%

0
1

ENG OIL
ENGINE OIL
TEMPERATURE

TEMP
C X 10

PRESS
PSI X 10

18
14

13
11
9

30−MINUTE LIMIT

135 − 150OC

CONTINUOUS

−50 − 135 C

CONTINUOUS

4
0
−4
1

ENGINE OIL
PRESSURE

4
3
2
1

700

20 − 100
PSI*

* 35 PSI MINIMUM AT 90% Ng AND ABOVE

LEGEND
RED
YELLOW
GREEN
DIGITAL
READOUT

AA8671_1A
SA

Figure 5-2. Instrument Markings (Sheet 1 of 2)

700

Change 10

5-5

TM 1-1520-237-10

% TRQ

ENGINE % TRQ
700

1 2
140

140

120

100

10−SECOND
TRANSIENT
DUAL−ENGINE
SINGLE−ENGINE

100% − 125%
110% − 135%

CONTINUOUS
SINGLE−ENGINE
ONLY

0% − 110%

CONTINUOUS
DUAL−ENGINE

0% − 100%

1 2

TGT
TEMP
O

TURBINE GAS
TEMPERATURE

C X 100

700

850 − 886OC

10−SECOND
TRANSIENT

7
6

850 C

START ABORT
LIMIT

5
4
2

775 − 850OC

30−MINUTE
LIMIT

0
1

538 − 775OC

NORMAL

TGT

XMSN
PRESS
PSI X 10

TEMP
C X 10

MAIN TRANSMISSION
OIL TEMPERATURE
UH60A

EH
10

PRECAUTIONARY

105 − 120OC

CONTINUOUS

−50 − 105 C

19
11
7
6

6
4
0
−4

MAIN TRANSMISSION
OIL PRESSURE
PRECAUTIONARY

65 − 130 PSI

CONTINUOUS

30 − 65 PSI

IDLE AND
TRANSIENT

20 − 30 PSI

5
4
3

AA8671_2C
SA

Figure 5-2. Instrument Markings (Sheet 2 of 2)

5-6

Change 10

TM 1-1520-237-10

% TRQ

ENGINE % TRQ
10−SECOND TRANSIENT
1

NOTE
HELICOPTERS PRIOR TO S / N 91−26354 THAT
ARE NOT EQUIPPED WITH IMPROVED MAIN
ROTOR FLIGHT CONTROLS ARE FURTHER
RESTRICTED ABOVE 80 KIAS TO DUAL−
ENGINE CONTINUOUS TORQUE LIMITS
AS INDICATED BY A PLACARD ON THE
INSTRUMENT PANEL. SEE FIGURE 5−4.

140

120

100

0
1

DUAL−ENGINE
ABOVE 80 KIAS
80 KIAS OR BELOW

100% − 144%
120% − 144%

SINGLE−ENGINE

135% − 144%

CONTINUOUS
SINGLE−ENGINE

0% − 135%

DUAL−ENGINE
ABOVE 80 KIAS
AT OR BELOW 80 KIAS

0% − 100%
0% − 120%

TGT
TEMP
C X 100

TURBINE GAS
TEMPERATURE
10−SECOND
TRANSIENT

903 − 949OC

2.5−MINUTES
TRANSIENT
(CONTINGENCY
POWER)

878 − 903OC

START ABORT LIMIT

851OC

10−MINUTE
LIMIT

851 − 878OC

30−MINUTE
LIMIT

810 − 851OC

NORMAL

538 − 810OC

9
8
7
6
5
4
2
1

TGT

LEGEND
RED
YELLOW
GREEN
DIGITAL READOUT

AA8672_1B
SA

Figure 5-3. Instrument Markings (Sheet 1 of 2)

701C

Change 10

5-7

TM 1-1520-237-10

ENG OIL

ENGINE OIL
TEMPERATURE

TEMP
O
C X 10

PRESS
PSI

170

135 − 150OC

CONTINUOUS

− 50 − 135OC

701C

120
100

30−MINUTE LIMIT

ENGINE OIL
PRESSURE

10
8

100 − 120 PSI

5−MINUTE
LIMIT

NORMAL OPERATION

26 − 100 PSI

IDLE

22 − 26 PSI

−4

XMSN
PRESS
PSI

TEMP
C X 10

190

MAIN TRANSMISSION
OIL TEMPERATURE

UH60L

PRECAUTIONARY
CONTINUOUS

110

− 50 − 105OC

PRECAUTIONARY

65 − 130 PSI

CONTINUOUS

30 − 65 PSI

IDLE AND
TRANSIENT

20 − 30 PSI

70
60

105 − 140OC

MAIN TRANSMISSION
OIL PRESSURE

0
−4

Ng
SPEED
% X 10

11
10

ENGINE Ng

9
8
7

4
1

10−SECOND
TRANSIENT

102% − 105%

30−MINUTE
LIMIT

99% − 102%

CONTINUOUS

0 − 99%

AA8672_2C
SA

Figure 5-3. Instrument Markings (Sheet 2 of 2)

5-8

Change 10

TM 1-1520-237-10

Section III POWER LIMITS
5.8 ENGINE LIMITATIONS.

tive start cycles. A 30-minute rest period is then required
before any additional starts.

5.8.1 Engine Power Limitations. 700 The limitations
which are presented in Figure 5-2, present absolute limitations, regardless of atmospheric conditions. For variations
in power available with temperature and pressure altitude,
refer to the TORQUE AVAILABLE charts in Chapter 7.

c. At ambient temperatures above 15° up to 52°C (59°
up to 126°F), two consecutive start cycles may be made. A
30-minute rest period is then required before any additional
start cycles.

5.8.2 Engine Power Limitations.

5.9 PNEUMATIC SOURCE INLET LIMITS.

701C

a. The limitations which are presented in Figure 5-3,
present absolute limitations regardless of atmospheric conditions. For variations in power available with temperature
and pressure altitude, refer to TORQUE AVAILABLE
charts in Chapter 7A.

The minimum ground-air source (pneumatic) required to
start the helicopter engines is 40 psig and 30 ppm at 149°C
(300°F). The maximum ground-air source to be applied to
the helicopter is 50 psig at 249°C (480°F), measured at the
external air connector on the fuselage.

b. Helicopters prior to S/N 91-26354 that are not
equipped with improved main rotor flight controls are further restricted above 80 KIAS to dual-engine continuous
torque limits as indicated by a placard on the instrument
panel . See Figure 5-4.

5.10 ENGINE START LIMITS.

5.8.3 Engine % RPM Limitations. Transient % RPM
1 or 2 operation in yellow range (101% to 105%) is not
recommended as good operating practice. However no
damage to either engine or drive train is incurred by operation within this range. Momentary transients above 107%
Np are authorized for use by maintenance test pilots during
autorotational rpm checks.
5.8.4 Engine Starter Limits.
a. The pneumatic starter is capable of making the number of consecutive start cycles listed below, when exposed
to the environmental conditions specified, with an interval
of at least 60 seconds between the completion of one cycle
and the beginning of the next cycle. A starting cycle is the
interval from start initiation and acceleration of the compressor, from zero rpm, to starter dropout. The 60-second
delay between start attempts applies when the first attempt
is aborted for any reason, and it applies regardless of the
duration of the first attempt. If motoring is required for an
emergency, the 60-second delay does not apply.
b. At ambient temperatures of 15°C (59°F) and below,
two consecutive start cycles may be made, followed by a
3-minute rest period, followed by two additional consecu-

CAUTION

Engine start attempts at or above a pressure altitude of 18,000 feet 701C , or
20,000 feet 700 could result in a Hot Start.
Crossbleed starts shall not be attempted unless the antiice light is off, and operating engine must be at 90% Ng
SPEED or above and rotor speed at 100% RPM R. When
attempting single-engine starts at pressure altitudes above
14,000 feet, press the start switch with the ENG POWER
CONT lever OFF, until the maximum motoring speed
(about 24%) is reached, before going to IDLE. Engine
starts using APU source may be attempted when within the
range of FAT and pressure altitude of Figure 5-5.
5.11 ENGINE OVERSPEED CHECK LIMITATIONS.
Engine overspeed check in flight is prohibited. Engine
overspeed checks, on the ground, are authorized by designated maintenance personnel only.
5.12 FUEL LIMITATIONS.
When using all fuel types, both fuel boost pumps shall
be on and operational, otherwise engine flameout may result.

Change 6

5-9

TM 1-1520-237-10

UH−60L DUAL−ENGINE TORQUE LIMITS − % TORQUE
−10

8
6
5

100

4
3
FAT, O C

100% TORQUE 2000 FT & BELOW
−20

−10

FAT, O C

PRESSURE ALT 1000 FT

−20

PRESSURE ALT 1000 FT

FAT, O C

HELICOPTERS PRIOR TO S / N 91−26354 NOT EQUIPPED WITH IMPROVED MAIN ROTOR FLIGHT CONTROLS.
AA1641A
SA

Figure 5-4. Dual-Engine Torque Limitations at Airspeeds Above 80 KIAS

5-10

701C

TM 1-1520-237-10

ENGINE START ENVELOPE

20
700 ALT LIMIT

EXAMPLE
WANTED
IF TWO−ENGINE START CAN BE
DONE AT 2900 FEET PRESSURE
ALTITUDE AND 16 OC

KNOWN
PRESSURE ALTITUDE = 2900 FEET
FREE−AIR TEMPERATURE = 16 OC

METHOD

PRESSURE ALTITUDE ~ FEET X 1000

18
701C ALT LIMIT

SINGLE ENGINE
START LIMIT

6
DUAL ENGINE
START LIMIT
4

ENTER CHART AT PRESSURE
ALTITUDE 2900 FEET
MOVE RIGHT TO INTERSECT
VERTICAL TEMPERATURE LINE.
IF LINES INTERSECT WITHIN DARK
SHADED AREA, TWO−ENGINE
START CAN BE DONE.

0
−60

−50

−40

−30

−20

−10

FREE−AIR TEMPERATURE ~ O C

AA0700A
SA

Figure 5-5. Engine Start Envelope

5-11

TM 1-1520-237-10

Section IV LOADING LIMITS
5.13 CENTER OF GRAVITY LIMITATIONS.
Center of gravity limits for the aircraft to which this
manual applies and instructions for computation of the center of gravity are contained in Chapter 6.

5. When operating at or above gross weights of 20,500
pounds, the seven lug wheel may experience lug failure
resulting in flying debris during ground handling and/or
unexpected tire failure. The fourteen lug wheel shall be
utilized when operating at or above gross weights of 20,500
pounds.

5.14 WEIGHT LIMITATIONS.
5.15 STOWAGE PROVISIONS.
AIRCRAFT

MAXIMUM WEIGHT

Maximum capacity for each storage compartment is 125
pounds.

UH-60A

20,250

EH-60A

20,250

UH-60A (see paragraph 1)

22,000

EH-60A (see paragraph 1)

22,000

UH-60L

22,000

UH-60A/L with seven lug
wheels (see paragraph 5)

20,500

UH-60L External lift mission (see paragraph 3)

23,500

For UH-60A aircraft, the maximum weight that may be
suspended from the cargo hook is limited to 8,000 pounds.
For UH-60L aircraft, the maximum weight that can be suspended from the cargo hook is 9,000 pounds.

ESSS aircraft on ferry
mission (see paragraph 2)

24,500

NOTE

1. UH-60A and EH-60A maximum gross weight can be
extended from 20,250 pounds to 22,000 pounds only when
wedge mounted pitot-static probes and either/or MWO 551520-237-50-58 or MWO 1-1520-237-50-73 are installed.

5.16 CABIN CEILING TIEDOWN FITTINGS.
The four cabin ceiling tiedown fittings have a limited
load capability of 4,000 pounds.
5.17 CARGO HOOK WEIGHT LIMITATION.

UH-60L aircraft prior to serial number 9226421, will require an entry into DA Form
2408-13-1 following the first mission carrying an external cargo hook load exceeding
8,000 pounds.

2. Airworthiness release required.
5.18 RESCUE HOIST WEIGHT LIMITATIONS.
3. External lift missions above 22,000 pounds can only
be flown with cargo hook loads above 8,000 pounds and up
to 9,000 pounds.
4. Maximum weight is further limited by cargo floor
maximum capacity of 300 pounds per square foot. Refer to
Chapter 6.

5-12

Change 9

The maximum weight that may be suspended from the
rescue hoist is 600 pounds.

TM 1-1520-237-10

Section V AIRSPEED LIMITS
5.19 AIRSPEED OPERATING LIMITS.
The airspeed operating limits charts (Figures 5-6, 5-7
and 5-8) define velocity never exceed (Vne) as a function
of altitude, temperature, and gross weight. The dashed lines
represent the Mach limited airspeeds due to compressibility
effects. Additional airspeed limits not shown on the charts
are:
a. Maximum airspeed with external cargo hook loads
greater than 8,000 pounds and a corresponding gross weight
greater than 22,000 pounds will vary due to the external
load physical configuration, but shall not exceed 120 KIAS.
b. Maximum airspeed for one engine inoperative is 130
KIAS.
c. Maximum airspeed for autorotation at a gross weight
of 16,825 pounds or less is 150 KIAS.

airspeed of 130 KIAS. With landing light extended, airspeed is limited to 180 KIAS.
(2) Searchlight. If use is required, the searchlight must
be extended prior to reaching a maximum forward airspeed
of 100 KIAS. With searchlight extended, airspeed is limited
to 180 KIAS.
i. VOL The maximum airspeed for autorotation shall
be limited to 100 KIAS.
j. Maximum airspeed with skis installed is 155 KIAS.
5.20 FLIGHT WITH CABIN DOOR(S)/WINDOW(S)
OPEN.
The following airspeed limitations are for operating the
helicopter in forward flight with the cabin doors/window
open:
a. Cabin doors.

d. Maximum airspeed for autorotation at a gross weight
of greater than 16,825 pounds is 130 KIAS.
e. Sideward/rearward flight limits. Hovering in winds
greater than 45 knots (35 knots with external ERFS) from
the sides or rear is prohibited. Sideward/rearward flight into
the wind, when combined with windspeed, shall not exceed
45 knots (35 knots with external ERFS).
f. SAS inoperative airspeed limits:
(1) One SAS inoperative - 170 KIAS.

(1) Cabin doors may be fully open up to 100 KIAS
with soundproofing installed aft of station 379.
(2) Cabin doors may be fully open up to 145 KIAS
with soundproofing removed aft of station 379 or with
soundproofing secured properly.
(3) The doors will not be intentionally moved from the
fully open or closed position in flight. The cabin doors may
be opened or closed during hovering flight. The cabin doors
must be closed or fully opened and latched before forward
flight. Should the door inadvertently open in flight, it may
be secured fully open or closed.

(2) Two SAS inoperative - 150 KIAS.
(3) Two SAS inoperative in IMC - 140 KIAS.
g. Hydraulic system inoperative limits:

b. Gunner’s window(s) may be fully open up to 170
KIAS.
c. Cockpit doors sliding windows will not be opened or
closed during flight except during hover.

(1) One hydraulic system inoperative - 170 KIAS.

d. Flight with cockpit door(s) removed is prohibited.

(2) Two hydraulic systems inoperative - 150 KIAS.

e. VOL Flight with cabin door(s) open is not authorized.

(3) Two hydraulic systems inoperative in IMC - 140
KIAS.
h. Searchlight and landing light airspeed limits.
(1) Landing light. If use is required, the landing light
must be extended prior to reaching a maximum forward

5.21 AIRSPEED LIMITATIONS FOLLOWING FAILURE OF THE AUTOMATIC STABILATOR CONTROL SYSTEM.
a. Manual control available. If the automatic stabilator
control system fails in flight and operation cannot be restored:

Change 9

5-13

TM 1-1520-237-10

AIRSPEED OPERATING LIMITATIONS
100% RPM R

AIRCRAFT WITHOUT ESSS INSTALLED
EXAMPLE
WANTED

−50

MAX IAS FOR VARIOUS
TEMPS, PRESSURE
ALTITUDE AND
GROSS WEIGHTS

20
0

40
0

200

20
30

FREE AIR TEMPERATURE ~ O C

00
60

−2

40
50

0
00
0
18
00
0
17
00
0
16
00
0
15
00
0

LE
SS

14
00
0

110

GROSS
WEIGHT
~ LBS

100

MAXIMUM INDICATED AIRSPEED (VNE) ~ KNOTS

90
21

ENTER FAT AT −20oC.
MOVE RIGHT TO
PRESSURE ALTITUDE
4,000 FEET.
MOVE DOWN TO
18,000 POUNDS
GROSS WEIGHT
OR MACH LIMIT
FAT WHICHEVER
IS ENCOUNTERED
FIRST, IN THIS
CASE 18,000
POUNDS IS
ENCOUNTERED
FIRST. MOVE LEFT
TO READ 186
KNOTS.

−10

0
00
14

METHOD

−20

TT
0
~~FF
00
12
DDEE
TUU
TTITI
AALL 00
RREE 000000
SSUU 11
EESS
PPRR
00
80

FAT = − 20oC
PRESSURE ALTITUDE
= 4,000 FEET.
GROSS WEIGHT
= 18,000 POUNDS.

−30

0
00
18

KNOWN

−40

120
−40oC
130
−30oC

−50OC
140
150
−20oC
160

170

−10oC

180
COMPRESSIBILITY
LIMITS ~ FAT
190
200
−2

DENSITY ALTITUDE ~ 1000 FEET

22
AA1250A
SA

Figure 5-6. Airspeed Operating Limits
5-14

TM 1-1520-237-10

AIRSPEED OPERATING LIMITATIONS
AIRCRAFT WITH EXTERNAL STORES
SUPPORT SYSTEM INSTALLED
100% RPM R
−50
−40
00
0

00
0

00
60

−10

0
00
14

0
FT
00
E~
12
UD
TIT
AL 0
RE 00
0
SU 1
ES
PR
00
80

−20

40
0

200

20
30

FREE AIR TEMPERATURE ~ O C

20
18

−30

−2

40
50

24
24 500
0
23 00
GH
00
T
0
=
22
0
21 00
20 000 LB
19 000
18 000
0
17 00
16 000
15 000
14 000
00
0

FAT
~ OC

100

110

−50

120

LIM
ITS

−40

130
−30

140
150

MAXIMUM INDICATED AIRSPEED ~ KNOTS

−20

160
170
180
−2

22
AA1251B

DENSITY ALTITUDE ~ 1000 FEET

Figure 5-7. Airspeed Operating Limits

5-15

TM 1-1520-237-10

AIRSPEED OPERATING LIMITATIONS
VOLCANO MINE DESPENSING SYSTEM
WITH CANISTERS
100% RPM R

AIRSPEED
OPERATING
LIMITS
VOLCANO

−60
−50
20

FREE AIR TEMPERATURE ~OC

−40

PRESSURE
ALTITUDE
~ 1000 FT

−30
16

−20

14
12

−10
10

0
8

6
4

20
2

30
0

−2

MAXIMUM INDICATED AIRSPEED (VNE) ~ KTS

60
70
GROSS
WEIGHT
~ 1000 LB

22
20

18
90

16
14

100

−50
110
120
−40

FAT
~ OC

130
T

IMI

C
MA

140
150
−2

20
AA9440

DENSITY ALTITUDE ~ 1000 FT

Figure 5-8. Airspeed Operating Limits

5-16

VOL

(Sheet 1 of 2)

TM 1-1520-237-10

AIRSPEED OPERATING LIMITATIONS
VOLCANO MINE DESPENSING SYSTEM
WITHOUT CANISTERS
100% RPM R

AIRSPEED
OPERATING
LIMITS
VOLCANO

−60
−50
20

FREE AIR TEMPERATURE ~OC

−40

PRESSURE
ALTITUDE
~ 1000 FT

−30
16

−20

14
12

−10
10

0
8

6
4

20
2

30
0

−2

MAXIMUM INDICATED AIRSPEED (VNE) ~ KTS

60
70
80
GROSS
WEIGHT
~ 1000 LB

22
20
18

100

16
−50

110

−40

FAT
~ OC

120
130
C
MA

140

150
160
−2

20
AA9441

DENSITY ALTITUDE ~ 1000 FT

Figure 5-8. Airspeed Operating Limits

VOL

(Sheet 2 of 2)

5-17

TM 1-1520-237-10

(1) The stabilator shall be set full down at speeds below 40 KIAS.
(2) The stabilator shall be set at zero degrees at speeds
above 40 KIAS.
(3) Autorotation airspeed shall be limited to 120 KIAS
at all gross weights.

5-18

b. Manual control not available. The placard airspeed
limits shall be observed as not-to-exceed speed (powered
flight and autorotation), except in no case shall the autorotation limit exceed 120 KIAS.

TM 1-1520-237-10

Section VI MANEUVERING LIMITS
5.22 PROHIBITED MANEUVERS.
a. Hovering turns greater than 30° per second are prohibited. Intentional maneuvers beyond attitudes of 630° in
pitch or over 60° in roll are prohibited.
b. Simultaneous moving of both ENG POWER CONT
levers to IDLE or OFF (throttle chop) in flight is prohibited.
c. Rearward ground taxi is prohibited.
5.23 RESTRICTED MANEUVERS.
5.23.1 Manual Operation of the Stabilator. Manual
operation of the stabilator in flight is prohibited except as
required by formal training and maintenance test flight requirements, or as alternate stabilator control in case the
AUTO mode malfunctions.
5.23.2 Downwind Hovering. Prolonged rearward flight
and downwind hovering are to be avoided to prevent accumulation of exhaust fumes in the helicopter and heat damage to windows on open cargo doors.
5.23.3 Maneuvering Limitations.
NOTE
Maneuvers entered from a low power setting
may result in transient droop of 5% RPM R
or greater.
a. The maneuvering limits of the helicopter, other than
as limited by other paragraphs within this section, are always defined by main rotor blade stall. Stall has not been
encountered in one G flight up to the airspeeds shown in
chart Figure 5-6 for aircraft without ESSS installed and
Figure 5-7 for aircraft with ESSS installed.

ducing the angle of bank. Maneuvering flight which results
in severe blade stall and significant increase in 4 per rev
vibration is prohibited.
5.23.3.1 High Speed Yaw Maneuver Limitation.
Above 80 KIAS avoid abrupt, full pedal inputs to prevent
excess tail rotor system loading.
5.23.3.2 Limitations for Maneuvering With Sling
Loads. Maneuvering limitations with a sling load is limited to a maximum of 30° angle of bank in forward flight
(Figure 5-10). Side flight is limited by bank angle and is
decreased as airspeed increases. Rearward flight with sling
load is limited to 35 knots.
5.23.3.3 Limitations for Maneuvering With Rescue
Hoist Loads. Maneuvering limitations with a rescue hoist
load is limited to maximum of 30° angle of bank in forward
flight (Figure 5-10). Side flight is limited by bank angle and
is decreased as airspeed is increased. Rearward flight with
hoist load is limited to 35 knots. Rate of descent is limited
to 1,000 feet-per-minute.
5.23.3.4 Bank Angle Limitation. Bank angles shall be
limited to 30° when a PRI SERVO PRESS caution light is
on.
5.24 LANDING GEAR LIMITATIONS.
Do not exceed a touchdown sink rate of 540 feet-perminute on level terrain and 360 feet-per-minute on slopes
with gross weights of up to 16,825 pounds; above 16,825
pounds gross weight 300 feet-per-minute on level terrain
and 180 feet-per-minute on slopes.
5.25 LANDING SPEED LIMITATIONS.
Maximum forward touchdown speed is limited to 60
knots ground speed on level terrain.
5.26 SLOPE LANDING LIMITATIONS.

b. The blade stall chart (Figure 5-9) while not an aircraft limitation, provides the level flight angle of bank at
which blade stall will begin to occur as a function of airspeed, gross weight, pressure altitude and temperature.
When operating near blade stall, any increase in airspeed,
load factor (bank angle), turbulence, or abrupt control inputs will increase the severity of the stall. Fully developed
stall will be accompanied by heavy four per rev vibration,
increasing torque, and loss of altitude. Recovery is always
accomplished by reducing the severity of the maneuver,
that is by reducing collective, reducing airspeed, and/or re-

The following slope limitations apply regardless of gross
weight or CG, with or without ESSS/ERFS.

CAUTION

When performing slope landings with External Extended Range Fuel System
Tanks, ensure tank to ground clearance.

Change 8

5-19

TM 1-1520-237-10

AIRSPEED FOR ONSET OF BLADE STALL
LEVEL FLIGHT 100% RPM R

EXAMPLE
WANTED
20

MAX RECOMMENDED
AIRSPEED FOR KNOWN
ANGLE OF BANK

PRESSURE ALTITUDE ~ 1000 FT

KNOWN
FAT = 20 OC
PRESSURE ALTITUDE
= 5,000 FEET.
GROSS WEIGHT
= 23,000 POUNDS
ANGLE OF BANK
= 20 DEGREES

METHOD
ENTER PRESSURE ALTITUDE
AT 5,000 FEET. MOVE
RIGHT TO 20 DEGREES FAT.
MOVE DOWN TO GROSS
WEIGHT 23,000 POUNDS. MOVE
LEFT TO 20 DEGREES ANGLE
OF BANK. MOVE VERTICALLY
DOWN TO READ INDICATED
AIRSPEED OF 108 KNOTS.

16
14
12
10
8

FAT OC

−60

−40

−20
0

60 40

0
10
20

VNE
ESSS

24.5
22

NOTE

WITH ESSS INSTALLED, REDUCE AIRSPEED
BY 6 KNOTS.

18
16
14

240

220

200

180

160

140

120

100

GROSS WEIGHT ~ 1000 LBS

VNE

ANGLE OF BANK ~ DEG

INDICATED AIRSPEED ~ KTS

AA1306A
SA

Figure 5-9. Airspeed for Onset of Blade Stall

5-20

TM 1-1520-237-10

NOTE
Because of the flat profile of the main transmission and forward location of both transmission oil pumps, transmission oil pressure
will drop during nose-up slope operations.
At slope angle of 10° an indicated oil pressure of 30 to 35 psi is normal, and at a 15°
slope angle a pressure in the range of 10 to
15 psi is normal, due to pitching of the helicopter.
a. 15° nose-up, right wheel up or left wheel upslope.
The slope limitations shall be further reduced by 2° for
every 5 knots of wind.
b. 6° nose downslope. Landing in downslope conditions with tail winds greater than 15 knots shall not be

conducted. A low-frequency oscillation may occur when
landing nose-down on a slope with the cyclic near the aft
stop.
c. The main gearbox may be operated up to 30 minutes
at a time with pressure fluctuations when the helicopter is
known to be at a nose-up attitude (i.e., slope landings or
hover with extreme aft CG).
d. When attempting a nose upslope landing at gross
weights in excess of 16,000 pounds with skis installed, the
parking brake may not hold the aircraft in position. The
pilot should be prepared to use the toe brakes.
e. Slope landings with skis installed are limited to 10°
nose up and right wheel or left wheel upslope.

Change 9

5-21

TM 1-1520-237-10

ANGLE OF BANK ~ O

SLING/RESCUE HOIST LOAD
MANEUVERING LIMITS
RESCUE
HOIST
LIMITS

ANGLE OF BANK LIMITS

SLING
LOAD
LIMITS

VNE FOR SLING LOAD
UP TO 8,000 POUNDS

VNE FOR EXTERNAL CARGO
SLING LOADS ABOVE 8,000
POUNDS AND CORRESPONDING
AIRCRAFT GROSS WEIGHTS IN
EXCESS OF 22,000 POUNDS
DUE TO EXTERNAL CARGO
HOOK LOAD

(SLING LOAD ENVELOPE)

0
40

100

120

140

KIAS

CROSSWIND
AND
SIDE FLIGHT
FORWARD FLIGHT

AA0668A
SA

Figure 5-10. Sling/Hoist Load Maneuvering Limitations

5-22

TM 1-1520-237-10

Section VII ENVIRONMENTAL RESTRICTIONS
5.27 FLIGHT IN INSTRUMENT METEOROLOGICAL
CONDITIONS (IMC).
This aircraft is qualified for operation in instrument meteorological conditions.
5.28 FLIGHT IN ICING CONDITIONS.
a. When the ambient air temperature is 4°C (39°F) or
below and visible liquid moisture is present, icing may occur. Icing severity is defined by the liquid water content
(LWC) of the outside air and measured in grams per cubic
meter (g/m3).
(1)
(2)
(3)
(4)

Trace
Light
Moderate
Heavy

:LWC
:LWC
:LWC
:LWC

0 to 0.25 g/m3
0.25 to 0.5 g/m3
0.5 to 1.0 g/m3
greater than 1.0 g/m3

b. Helicopters with the following equipment installed
and operational are permitted to fly into trace or light icing
conditions. Flight into light icing is not recommended without the blade deice kit. Flight into moderate icing shall
comply with paragraph 5.28 c.

of extreme low power requirements such as high rate of
descent (1900 fpm or greater), or ground operation below
100% RPM R, during icing conditions. The cabin heating
system should be turned off before initiating a high rate of
descent.
5.30 BACKUP HYDRAULIC PUMP HOT WEATHER
LIMITATIONS.
During prolonged ground operation of the backup pump
using MIL-H-83282 or MIL-H-5606 with the rotor system
static, the backup pump is limited to the following
temperature/time/cooldown limits because of hydraulic
fluid overheating.

FAT °C (°F)

Operating Time
(Minutes)

Cooldown Time
(Pump
Off)
(Minutes)

-54° - 32°
(-65° - 90°)
33° - 38°
(91° - 100°)
39° - 52°
(102° - 126°)

Unlimited

(1) Windshield Anti-ice.

5.31 APU OPERATING LIMITATIONS.

(2) Pitot Heat.

(4) Engine Inlet Anti-ice Modulating Valve.

To prevent APU overheating, APU operation at ambient
temperature of 43°C (109°F) and above with engine and
rotor operating, is limited to 30 minutes. With engine and
rotor not operating, the APU may be operated continuously
up to an ambient temperature of 51°C (124°F).

(5) Insulated Ambient Air Sensing Tube.

5.32 WINDSHIELD ANTI-ICE LIMITATIONS.

(3) Engine Anti-ice.

c. For flight into moderate icing conditions, all equipment in paragraph 5.28 b. and blade deice kit must be installed and operational. Flight into heavy or severe icing is
prohibited.

Windshield anti-ice check shall not be done when FAT
is over 27°C (80°F).
5.33 TURBULENCE AND THUNDERSTORM OPERATION.

d. Helicopters equipped with blade erosion kit are prohibited from flight into icing conditions.

a. Intentional flight into severe turbulence is prohibited.

5.29 ENGINE AND ENGINE INLET ANTI-ICE LIMITATIONS.

b. Intentional flight into thunderstorms is prohibited.

At engine power levels of 10% TRQ per engine and
below, full anti-ice capability cannot be provided, due to
engine bleed limitations. Avoid operation under conditions

c. Intentional flight into turbulence with a sling load attached and an inoperative collective pitch control friction is
prohibited.

Change 9

5-23

TM 1-1520-237-10

Section VIII OTHER LIMITATIONS
5.34 EXTERNAL EXTENDED RANGE FUEL SYSTEM KIT CONFIGURATIONS. ES

c. VOL Jettisoning, if necessary, shall be accomplished at airspeeds not to exceed 110 KIAS and rates of
descent not to exceed 500 fpm.

NOTE
5.36 ES USE OF M60D GUN(S) WITH ERFS KIT INSTALLED.

Flight with 450-gallon ERFS tanks is prohibited unless operating under an Airworthiness Release from U. S. Army Aviation and
Missile Command.
The ERFS kit shall only be utilized in the following
approved configurations:
a. A 230-gallon tank installed on each inboard vertical
stores pylon.

Use of the M60D gun(s) is prohibited when external
ERFS tanks are installed on the outboard vertical stores
pylons, unless the external ERFS pintle mount stop is installed. Use of the M60D gun(s) is prohibited when external tanks are installed on the inboard vertical stores pylon.
5.37 GUST LOCK LIMITATIONS.

WARNING

b. A 230-gallon tank installed on each outboard vertical
stores pylon.

Before engine operations can be performed with the gust lock engaged, all
main rotor tie downs shall be removed.

c. Four 230-gallon tanks installed, one on each inboard
and each outboard vertical stores pylon.

a. Dual-engine operation with gust lock engaged is prohibited.

5.35 JETTISON LIMITS.
a. ES The jettisoning of fuel tanks in other than an
emergency is prohibited.
b. ES The recommended external fuel tank jettison envelope is shown in Table 5-1.

b. Single-engine operation with gust lock engaged will
be performed by authorized pilot(s) at IDLE only.
c. Gust lock shall not be disengaged with engine running.

Table 5-1. Recommended Emergency External Fuel Tank Jettison Envelope
RECOMMENDED EMERGENCY JETTISON ENVELOPE
AIRSPEED KIAS
0 TO 120

120 TO Vh

SLIP INDICATOR DISPLACED NO MORE THAN
ONE BALL WIDTH LEFT OR RIGHT

NO
SIDESLIP
BALL
CENTERED

LEVEL FLIGHT

AIRSPEED KIAS
DESCENT

*JETTISON
BELOW
80 KIAS
NOT
RECOMMENDED

100

110

120

1000

875

750

625

500

MAX RATE OF DESCENT
FT/MIN

*JETTISON
ABOVE
120 KIAS
NOT
RECOMMENDED

*Not recommended because safe jettison at these conditions has not been verified by tests.

5-24

Change 10

TM 1-1520-237-10

5.38 MAINTENANCE
(MOC).

OPERATIONAL

CHECKS

Maintenance operational checks (MOC) will be accomplished in accordance with TM 1-1500-328-23.

b. Use of GPS landing mode of CIS is prohibited under
IMC.
5.41 USE OF SKIS.
Water buckets shall not be used when skis are installed.

5.39 USE OF AN/ARC-220() HF RADIO.
If installation of the AN/ARC-220() HF radio is not in
accordance with MWO 1-1520-237-50-76, an airworthiness
release from U. S. Army Aviation and Missile Command is
required.
5.40 USE OF AN/ASN-128B DOPPLER/GPS RADIO.
a. The AN/ASN-128B shall not be used as the primary
source of navigation information for Instrument Flight Rule
(IFR) operations in controlled airspace.

Change 10

5-25/(5-26 Blank)

TM 1-1520-237-10

CHAPTER 6
WEIGHT/BALANCE AND LOADING
Section I GENERAL
6.1 INTRODUCTION.
This chapter contains instructions and data to compute
any combination of weight and balance for this helicopter,
if basic weight and moment are known.
6.2 CLASS.
Army helicopters defined in this manual are in Class 1
and Class 2 UH . Additional directives governing
weight and balance of Class 1 and Class 2 aircraft forms
and records are contained in AR 95 series, TM 55-1500342-23, and DA PAM 738-751.

6.3 HELICOPTER COMPARTMENT AND STATION
DIAGRAM.
Figure 6-1 shows the reference datum line that is 341.2
inches forward of the centroid of the main rotor, the fuselage stations, waterlines and buttlines. The fuselage is divided into compartments A through F. The equipment in
each compartment is listed on DD Form 365-1 (Chart A) in
the individual aircraft weight and balance file.

Change 6

6-1

130
110
90
70
50
30
10
10
30
50
70
90
110
130

BUTT LINES

TM 1-1520-237-10

120
100
80
60
40
20

BL
0

BL 0

20
40
60
80
100
120

50 100

150

200

250

300

350

400

450

500

550

600

650

700

750

800

STATIONS

COMPARTMENTS

F
STA
732

STA
341.2

350
WL
324.7

WL
315

WATER LINES

300

250
WL
215

COCKPIT
FLOOR

200
STATIC
GROUND LINE

WL
206.7
STA
204

150
CABIN FLOOR
0

STA
162

STA
343
STA
247
STA
270

STA
315.5

STA
485

STA
398

STA
370.5

STA
644.6

STA
443.5

STA
763.5

AA0374
SA

Figure 6-1. Helicopter Compartment and Station Diagram

6-2

TM 1-1520-237-10

Section II WEIGHT AND BALANCE
6.4 SCOPE.
This section provides appropriate information required
for the computation of weight and balance for loading an
individual helicopter. The forms currently in use are the
DD Form 365 series. The crewmember has available the
current basic weight and moment which is obtained from
DD Form 365-3 (Chart C) for the individual helicopter.
This chapter contains weight and balance definitions; explanation of, and figures showing weights and moments of
variable load items.
6.5 WEIGHT DEFINITIONS.
a. Basic Weight. Basic weight of an aircraft is that
weight which includes all hydraulic systems and oil systems full, trapped and unusable fuel, and all fixed equipment, to which it is only necessary to add the crew, fuel,
cargo, and ammunition (if carried) to determine the gross
weight for the aircraft. The basic weight varies with structural modifications and changes of fixed aircraft equipment.
b. Operating Weight. Operating weight includes the basic weight plus aircrew, the aircrew’s baggage, and emergency and other equipment that may be required. Operating
weight does not include the weight of fuel, ammunition,
cargo, passengers or external auxiliary fuel tanks if such
tanks are to be disposed of during flight.
c. Gross Weight. Gross weight is the total weight of an
aircraft and its contents.
6.6 BALANCE DEFINITIONS.
6.6.1 Horizontal Reference Datum. The horizontal
reference datum line is an imaginary vertical plane at or
forward of the nose of the helicopter from which all horizontal distances are measured for balance purposes. Diagrams of each helicopter show this reference datum line as
balance station zero.
6.6.2 Arm. Arm, for balance purposes, is the horizontal
distance in inches from the reference datum line to the CG
of the item. Arm may be determined from the helicopter
diagram in Figure 6-1.
6.6.3 Moment. Moment is the weight of an item multiplied by its arm. Moment divided by a constant is generally
used to simplify balance calculations by reducing the number of digits. For this helicopter, moment/1000 has been
used.

6.6.4 Average Arm. Average arm is the arm obtained
by adding the weights and moments of a number of items,
and dividing the total moment by the total weight.
6.6.5 Basic Moment. Basic moment is the sum of the
moments for all items making up the basic weight. When
using data from an actual weighing of a helicopter, the
basic moment is the total of the basic helicopter with respect to the reference datum. Basic moment used for computing DD Form 365-4 is the last entry on DD Form 365-3
for the specific helicopter. Cargo Hook Moments and Rescue Hoist Moments are shown in Figures 6-7 and 6-8, respectively.
6.6.6 Center of Gravity (CG). Center of gravity is the
point about which a helicopter would balance if suspended.
Its distance from the reference datum line is found by dividing the total moment by the gross weight of the helicopter.
6.6.7 CG Limits. CG limits (Figures 6-13 and 6-14) defines the permissible range for CG stations. The CG of the
loaded helicopter must be within these limits at takeoff, in
the air, and on landing.
6.7 DD FORM 365-3 (CHART C) WEIGHT AND BALANCE RECORDS.
DD Form 365-3 (Chart C) is a continuous history of the
basic weight, moment, and balance, resulting from structural and equipment changes in service. At all times the last
weight, moment/constant, is considered the current weight
and balance status of the basic helicopter.
6.8 LOADING DATA.
The loading data in this chapter is intended to provide
information necessary to work a loading problem for the
helicopter. From the figures, weight and moment are obtained for all variable load items and are added arithmetically to the current basic weight and moment from DD
Form 365-3 (Chart C) to obtain the gross weight and moment. If the helicopter is loaded within the forward and aft
CG limits, the moment figure will fall numerically between
the limiting moments. The effect on the CG of the expenditures in flight of such items as fuel and cargo may be
checked by subtracting the weights and moments of such
items from the takeoff gross weight and moment, and
checking the new moment, with the CG limits chart. This
check should be made to determine whether or not the CG
will remain within limits during the entire flight.

Change 8

6-3

TM 1-1520-237-10

6.9 DD FORM 365-4 (FORM F).
There are two versions of DD Form 365-4. Refer to TM
55-1500-342-23 for completing the form.

6-4

TM 1-1520-237-10

Section III FUEL/OIL
6.10 FUEL MOMENTS.

CAUTION

Fuel transfer sequence must be carefully
planned and executed in order to maintain CG within limits.
When operating with a light cabin load or
no load, it may be necessary to adjust fuel
load to remain within aft CG limits. Fuel
loading is likely to be more restricted on
those aircraft with the HIRSS installed.
For a given weight of fuel there is only a very small
variation in fuel moment with change in fuel specific
weight. Fuel moments should be determined from the line
on Figure 6-2 which represents the specific weight closest
to that of the fuel being used. The full tank usable fuel
weight will vary depending upon fuel specific weight. The
aircraft fuel gage system was designed for use with JP-4,

but does tend to compensate for other fuels and provide
acceptable readings. When possible the weight of fuel onboard should be determined by direct reference to the aircraft fuel gages. The following information is provided to
show the general range of fuel specific weights to be expected. Specific weight of fuel will vary depending on fuel
temperature. Specific weight will decrease as fuel temperature rises and increases as fuel temperature decreases at the
rate of approximately 0.1 lb/gal for each 15°C change. Specific weight may also vary between lots of the same type
fuel at the same temperature by as much as 0.5 lb/gal. The
following approximate fuel weights at 15°C may be used
for most mission planning:

Fuel Type

Specific Weight

JP-4

6.5 lb/gal.

JP-5

6.8 lb/gal.

JP-8

6.7 lb/gal.

Jet A

6.8 lb/gal.

Jet B

6.3 lb/gal.

Change 8

6-5

TM 1-1520-237-10

FUEL MOMENTS

EXAMPLE
WANTED
FUEL MOMENT

KNOWN

ITEM

STA

FUEL QUANTITY
MAIN 1700 POUNDS

230−GALLON TANK (IB OR OB)

321

150

450−GALLON TANK (IB)

316

234

MOM/1000

WEIGHT LBS

METHOD
FOR MAIN TANK ENTER
AT 1700 POUNDS AND
MOVE RIGHT TO MAIN LINE.
MOVE DOWN READ
MOMENT / 1000 = 710

ARM = 314.6 = 450−GALLON TANK
319.9 = 230−GALLON TANK

GALLONS
JP−5
ARM = 420.75 JP−4
450

3000

450
400
400
2500
350

FUEL WEIGHT (POUNDS)

350
300

2000

300

250
1500

200
1000

200

150

100

ARM = 319.9
OUTBOARD TANK

500

0
0

200

400

600

FUEL MOMENT/1000

Figure 6-2. Fuel Moments

6-6

250

INBOARD TANK

800

1000
AA0380E
SA

TM 1-1520-237-10

Section IV PERSONNEL
6.11 PERSONNEL MOMENTS.
When aircraft are operated at critical gross weights, the
exact weight of each individual occupant plus equipment
should be used. Personnel moments data is shown on Figure 6-3. If weighing facilities are not available, or if the
tactical situation dictates otherwise, loads shall be computed as follows:
a. Combat equipped soldiers: 240 pounds per individual.

6.12 MEDEVAC KIT PERSONNEL MOMENTS.
a. Litter moments are in Figure 6-4.
b. Medevac system (excluding litters) weight and moments are included in the helicopter basic weight and moments Form 365-3 when installed.
c. Litter weight is estimated to 25 pounds which includes litter, splints, and blankets.

b. Combat equipped paratroopers: 260 pounds per individual.

d. Medical attendant’s average weight is 200 pounds.

c. Crew and passengers with no equipment: compute
weight according to each individual’s estimate.

e. Medical equipment and supplies should be stored per
unit loading plan and considered in weight and balance
computations.

6-7

TM 1-1520-237-10

PERSONNEL MOMENTS
A

BL
40.0

CREWCHIEF /
GUNNER

PILOT

BL
20.0

BL
0

BL
0
BL
10.0

COPILOT

BL
20.0
BL
30.0
BL
40.0

STA
227.1

STA
204

STA
162.0

STA
247.0

STA
262.0

STA
282.0

ROW
1

ROW
2

STA
288.0

STA
320.7

STA
339.8

ROW
3

ROW
4

STA
343.0

STA
387.2

STA
398.0

ROW
5

SEAT WEIGHT − AND MOMENT TABLE*
ITEM
CREWCHIEF / GUNNER (2)
TROOPS (3)
TROOPS (3)
TROOPS (4)

ROW

WEIGHT

MOM / 1000

2
3
4
5

43
48
48
63

12
15
16
25

202

TOTAL−12 SEATS

EXAMPLE
WANTED:
PESONNEL MOMENTS

ALTERNATE SEATING (BROKEN LINES)
FORWARD TROOP
SEAT (NO SEAT AUTHORIZED IN
THIS POSITION)

*SEAT WEIGHT AND MOMENTS
SHOULD BE INCLUDED ON CHART C

KNOWN:

REAR FACING TROOP
SEAT (1)

2 PERSONNEL IN ROW 3
TOTAL WEIGHT 480 POUNDS

REAR FACING TROOP
SEAT (1)

METHOD:

234

TOTAL−14 SEATS

ENTER WEIGHT AT 480
POUNDS−MOVE RIGHT
TO ROW 3.
MOVE DOWN. READ
MOMENT / 1000=154

AA0669_1B
SA

Figure 6-3. Personnel Moments (Troop Configuration) (Sheet 1 of 3)

6-8

Change 10

TM 1-1520-237-10

PERSONNEL MOMENTS

M
RO
W

800

900

AR
M

38
7.
2

1000

20
=3
M
AR

600

2
W

400

227

500

PERSONNEL WEIGHT ~ POUNDS

700

OP
ILO

300

PIL

−C

200

100

0
0

100

150

200

250

300

350

400

MOMENT/1000

DATA BASIS:

AA0669_2A

CALCULATED

Figure 6-3. Personnel Moments (Troop Configuration) (Sheet 2 of 3)

Change 10

6-9

TM 1-1520-237-10

BL
40.0

ECM OPERATOR

PILOT

BL
20.0

OBSERVER

BL
0

COPILOT
BL
20.0

DF
OPERATOR

STA
162.0

STA
204

STA
227.1

STA
247.0

STA
288.0

* ITEM
OBSERVER SEAT
TOTAL − 1 SEAT

STA
324.5

STA
328.25

STA
343.0

STA
356.0

STA

WEIGHT

MOM / 1000

356.0

STA
398.0

* SEAT WEIGHT AND MOMENTS SHOULD BE INCLUDED ON CHART C.

STA
227.1

METHOD
ENTER WEIGHT AT 210
POUNDS − MOVE RIGHT
TO OBSERVER ARC (STA 356.0)
MOVE DOWN READ
MOMENT / 1000 = 75

T
LO
PI
200

100

DATA BASIS: CALCULATED

OBSERVER
STA 356.0

PERSONNEL AT STA 356
OBSERVER − 210 POUNDS

DF
OPERATOR
STA 328.25

300

T−
C

KNOWN

400

WANTED
PERSONNEL MOMENTS

ECM
OPERATOR
STA 324.5

EXAMPLE

PERSONNEL WEIGHT ~ POUNDS

500

100

125

MOMENT/1000
AA0669_3B
SA

Figure 6-3. Personnel Moments (EH Configuration) (Sheet 3 of 3)

6-10

TM 1-1520-237-10

LITTER MOMENTS

CREW CHIEF

PILOT

BL
0

COPILOT

MEDICAL ATTENDANT

STA
162.0

STA
204.0

STA
227.1

STA
247.0

STA
271.0

ROW
6

STA
288.0

STA
343.0

STA
343.6

ROW
7

CENTROID

STA
398.0

EXAMPLE
WANTED
LITTER MOMENTS

KNOWN
LITTER WEIGHT
= 265 POUNDS

METHOD
ENTER WEIGHT AT
265 POUNDS − MOVE
RIGHT TO LITTER
ROW 7
MOVE DOWN. READ
MOMENT / 1000 = 91

AA0378_1B
SA

Figure 6-4. Personnel Moments (Medevac Configuration) (Sheet 1 of 2)
Change 10

6-11

TM 1-1520-237-10

LITTER MOMENTS
ARM = 343.6

1100

1000

900

7
PA
TI

−R

700

600

T−

500

LITTER PATIENT WEIGHT POUNDS

800

DIC

400

300

200

ARM = 271.0

100

0
0

100

150

200

250

300

350

400

MOMENT/1000
DATA BASIS:CALCULATED

Figure 6-4. Personnel Moments (Medevac Configuration) (Sheet 2 of 2)
6-12

AA0378_2A
SA

TM 1-1520-237-10

Section V MISSION EQUIPMENT
6.13 ARMAMENT LOADING DATA MOMENTS.
Armament consists of two M60D machineguns, ammunition, and grenades. Various loads of ammunition are presented in Figure 6-5. When determining the moments for a
given ammo load not shown on the chart, go to the nearest
load shown. VOL Volcano mine moments are presented
in Figure 6-6.

6.14 EH-60A HELICOPTERS WITHOUT MISSION
EQUIPMENT.
When operating without EH-60 mission equipment
or with a light cabin load or no cabin load, it may be necessary to limit fuel load to remain within aft CG limits.
EH

6-13

TM 1-1520-237-10

ARMAMENT LOADING DATA
STA
279.8

AMMUNITION TABLE
LIVE
ROUNDS

LIVE AMMO (7.62 MM)
ARM − 247.0

AMMUNITION
BOX
EJECTION
BAG

WEIGHT − LB MOM / 1000
7
13
20
26
32
39
46
52

100
200
300
400
500
600
700
800

FIRING
POSITION
RH GUN

AMMO

2
3
5
6
8
10
11
13

GRENADES

ARM − 279.8
7
13
20
26

100
200
300
400

BL
0

2
4
5
7

CHAFF
CHAFF CARTRIDGE MI. 30 RDS
ARM − 505.0
(SINGLE CHAFF WEIGHT 0.33 LB)
WEIGHT − LB

MOM / 1000

FLARE EH

AMMO

STA
162

FLARE DISPENSED M130, 30 RDS
ARM − 525.0
(SINGLE FLARE WEIGHT 0.43 LB)
WEIGHT − LB

MOM / 1000

STOWED
POSITION
LH GUN

STA
247

GRENADES

(GUN STOWED AND FIRING POSITIONS ARE SAME EACH SIDE)

GRENADE TABLE

M60D TABLE
MOM / 1000

STOWED
GRENADE AN−M8
ARM − 251.0

QUANTITY

GRENADE M18
ARM − 251.0

WEIGHT − LB

MOM / 1000

WEIGHT − LB

MOM / 1000

3
6
9
12
15
18

1
2
2
3
4
5

2
5
7
10
12
14

1
1
2
2
3
4

2
4
6
8
10
12

DATA BASIS:

ITEM

WEIGHT
STOWED

FIRING
POSITION

M60D (2)
EJECTION BAG (2)
AMMO BOX (2)
STORAGE BOX (2)
SUPPORT (2)
BIPOD (2)

45.4
9.0
3.4
2.6
20.2
4.0

12
2
1
1
5
1

13
3
1
1
6
1

TOTAL

84.6

25
AA0375B

CALCULATED

Figure 6-5. Armament Loading Data Moments
6-14

STA
308

TM 1-1520-237-10

COLUMN
1

3
2

1
R
O
W
4

3
2

CANISTER COLUMN REFERENCE

AA9415
SA

Figure 6-6. Volcano Mine Moments

VOL

(Sheet 1 of 2)

6-15

TM 1-1520-237-10

RACK WEIGHTS (PER RACK)
NO CANISTERS
Weight (lb)

Arm

226

331.5

Moment/
1000
74.9

EMPTY CANISTERS (40)

FULL CANISTERS (40)

Weight (lb)

Arm

Moment/
1000

Weight (lb)

Arm

Moment/
1000

434

331.5

143.9

1450

331.5

480.7

WEIGHT (LB) QUANTITY
PER SYSTEM
Rack Without Canisters

TOTAL
WEIGHT (LB)

ARM MOMENT/1000

226

904

331.5

299.7

Empty

5.2

160

832

331.5

275.8

Full

30.6

160

4896

331.5

1623.0

Side Panels

236

472

322.8

152.4

DCU with Pallet

300

26.1

Cabling/ICP/Fairing
/Cable Tubes

264.4

14.3

Total System Full Canisters:

6413

329.7

2114.4

Total System Empty Canisters

2349

326.7

767.4

Total System No Canisters

1517

324.1

491.6

Canisters:

UNIT CANISTER LOADING
COLUMN EMPTY
CANISTER
WEIGHT

FULL
CANISTER
WEIGHT

ARM EMPTY
CANISTER
MOMENT/1000

5.2

30.6

306.3

1.6

9.4

5.2

30.6

311.8

1.6

9.5

5.2

30.6

317.3

1.6

9.7

5.2

30.6

322.8

1.7

9.9

5.2

30.6

328.3

1.7

10.0

5.2

30.6

333.8

1.7

10.2

5.2

30.6

339.3

1.8

10.4

5.2

30.6

344.8

1.8

10.6

5.2

30.6

350.3

1.8

10.7

5.2

30.6

355.8

1.9

10.9

Figure 6-6. Volcano Mine Moments
6-16

FULL
CANISTER
MOMENT/1000

Change 3

VOL

(Sheet 2 of 2)

TM 1-1520-237-10

CARGO MOMENTS − CARGO HOOK
ARM = 352.6

9000

8000

EXAMPLE
7000

WANTED
MOMENT OF CARGO
ON CARGO HOOK

6000

CARGO = 5600 POUNDS

METHOD
ENTER WEIGHT AT
5600 POUNDS. MOVE
RIGHT TO LINE. MOVE
DOWN AND READ
MOMENT / 1000
= 1975

WEIGHT ~ POUNDS

KNOWN

5000

4000

3000

2000

1000

0
0

1000

2000

CARGO HOOK MOMENTS/1000

3000

4000
AA8802
SA

Figure 6-7. Cargo Hook Moments

6-17

TM 1-1520-237-10

RESCUE HOIST MOMENTS
EXAMPLE

ARM = 367.5

600

WANTED
500
MOMENT OF RESCUE
HOIST LOAD

KNOWN

METHOD
ENTER WEIGHT AT
380 POUNDS − MOVE
RIGHT TO LINE. MOVE
DOWN. READ MOMENT /
1000 = 140

WEIGHT ~ POUNDS

400
RESCUE HOIST LOAD
= 380 POUNDS

300

200

100

0
0

DATA BASIS:

CALCULATED

100

MOMENT / 1000

Figure 6-8. Rescue Hoist Moments

6-18

150

200

250
AA0377
SA

TM 1-1520-237-10

Section VI CARGO LOADING
6.15 CABIN DIMENSIONS.

c. The helicopter’s center of gravity.

Refer to Figure 6-9 for dimensions. For loading, and
weight and balance purposes, the helicopter fuselage is divided into six compartments, A through F, three of which
are in the cabin, C, D, and E. There are 17 tiedown fittings
rated at 5,000 pounds each. Cargo carrier restraint rings are
at stations 308 and 379, to cover the 71 inches of longitudinal space. Cargo tiedown devices are stored in the equipment stowage space of compartment F.

d. Floor loads for each item of cargo.
e. Any shoring that may be required.
f. When required, the location of the center of gravity of
an individual item of cargo.
6.20.2 Cargo Center of Gravity Planning. The detail
planning procedure consists of four steps, as follows:

6.16 CABIN DOORS.
Cabin doors are at the rear of the cargo compartment on
each side of the fuselage. The door openings are 54.5 inches
high and 69 inches wide; maximum package sizes accommodated by the openings are 54 inches high by 68 inches
wide and are shown on Figure 6-10.

a. Determine ALLOWABLE LOAD from LIMITATIONS section of DD Form 365-4.

6.17 MAXIMUM CARGO SIZE DIAGRAM FOR
LOADING THROUGH CABIN DOORS.

b. Plan the location in the helicopter for the individual
items of cargo. Since the CG of the load is determined by
the station method, then specific locations must be assigned
to each item of cargo.

Figure 6-10 shows the largest size of cargo of various
shapes that can be loaded into the cabin through the cabin
doors.

c. Determine the CG of the cargo load as planned. Regardless of the quantity, type, or size of cargo, use the
station method.

6.18 TIEDOWN
RINGS.

FITTINGS

AND

RESTRAINT

The 17 tiedown fittings (Figure 6-11) installed on the
cargo floor can restrain a 5,000-pound load in any direction. All tiedown fittings incorporate studs that are used to
install the troop seats. Eight net restraint rings in the cargo
compartment prevent cargo from hitting the bulkhead at
station 398, or entering the crew area. Each restraint ring is
rated at 3500-pound capacity in any direction.
6.19 EQUIPMENT STOWAGE COMPARTMENTS.
Equipment stowage compartment moments are shown in
Figure 6-12.
6.20 EQUIPMENT LOADING AND UNLOADING.
6.20.1 Data Prior to Loading. The following data
should be assembled or gathered by the loading crew before loading (Refer to FM 55-450-2, Army Helicopter Internal Load Operations):
a. Weight of the individual items of cargo.
b. Overall dimensions of each item of cargo (in inches).

d. Determine the CG of the fully-loaded helicopter from
Figures 6-13 and 6-14, and if the CG of the helicopter falls
within allowable limits. If it does, the cargo can be loaded.
If not, the planned location of the individual items must be
changed until an acceptable loading plan is obtained. When
cargo loads consists of more than one item, the heavier
items of cargo should be placed so that their CG is about in
the center of the cabin, and the lighter items of cargo are
forward and rear of them.
6.20.3 Restraint Criteria. The amount of restraint that
must be used to keep the cargo from moving in any direction is called the 9restraint criteria9 and is usually expressed
in units of the force of gravity, of Gs. Following are the
units of the force of gravity or Gs needed to restrain cargo
in four directions:

Cargo
Forward

12 Gs

Rear

3 Gs

Lateral

8 Gs

Vertical

3 Gs (Up)
3 Gs (Down)

6-19

TM 1-1520-237-10

BL
34.5

BL
0

WL
261.0

FRAMES
69 INCHES

WL
234.0

85 INCHES
84 INCHES
72 INCHES

DOOR

WL
206.75

BL
36.0

STA 279.0
LOOKING TO THE REAR

BL
36.0

151 INCHES
STA
266.0

BL
29.0

BL
36.0

72 INCHES
AT FLOOR
LEVEL

BL
0

WL
206.75

BL
0

84 INCHES AT
CABIN DOORS
BL
10.0
BL
36.0

MR
CL
STA
247.0

STA
256.5

54.25
INCHES
52
INCHES

STA
341.2

STA
398.0

52
INCHES

DRIP
PAN

53.5
INCHES
68
INCHES

54.25
INCHES

CABIN
DOOR

CABIN AND DOOR DIMENSIONS

Figure 6-9. Cabin Dimensions

6-20

AA0379
SA

TM 1-1520-237-10

TIEDOWN RING
DOORWAY

RIGHT SIDE SHOWN (2 PLACES)
LEFT SIDE SAME (2 SHOWN)

MAXIMUM PACKAGE SIZE TABLE
CABIN DOORS

HEIGHT − INCHES

WIDTH
INCHES

50 &
UNDER

MAXIMUM LENGTH − INCHES

102

101

100

54"

68"

CABIN DOOR − BOTH SIDES

NOTE
IF GUNNERS AREA NOT USED,
LENGTHS ARE APPROXIMATELY 90%
OF TABLE VALUES.

AA0670
SA

Figure 6-10. Maximum Package Size for Cargo Door
6-21

TM 1-1520-237-10

WL
261.0

CARGO
RESTRAINT
NET RING

TOP OF
CABIN
FLOOR

3500 POUND
CAPACITY
EACH

WL
240.0

WL
206.75

STA 308.0
LOOKING TO THE FRONT

BL
0.0

STA 379.0
LOOKING TO THE FRONT

CARGO NETTING
EQUIPMENT STOWAGE
COMPARTMENTS (FORCE
RESTRAINT 1000 POUNDS
EACH)

STA 402.19 − LOOKING TO THE REAR

BL
0

BL
36.5

35 30 25 20 15 10 5
STA
247

BL
36.5

5 10 15 20 25 30 35

250
COMPARTMENT C

STA
288

255
260

TIEDOWN FITTING
5000 POUNDS CAPACITY

265
270

280

MAXIMUM
COMPARTMENT
CAPACITY
IN POUNDS

FLOOR CAPACITY
POUNDS PER
SQUARE FOOT

285

5460

300

8370

300

275

290
295
300
COMPARTMENT D

305
310
315
320
325
330
335
340

STA
343

345
350
COMPARTMENT E

355
360
365
370
375
380
385
390

STA
398

395
AA0671_1A
SA

Figure 6-11. Cargo Tiedown Arrangement
6-22

TM 1-1520-237-10

STOWAGE COMPARTMENT MOMENTS
STOWED SEAT TABLE
ITEM

ROW

WEIGHT

MOMENT / 1000

2
3
4
5

43
48
48
63

18
20
20
27

202

16
16
16

7
7
7

250

106

CREWCHIEF / GUNNER (2)
TROOPS (3)
TROOPS (3)
TROOPS (4)
TOTAL−12 SEATS

1
2
4

ALTERNATE (1)
ALTERNATE (1)
ALTERNATE (1)
TOTAL−15 SEATS

BL
32.9

BL
10.0

EXAMPLE

BL
10.0

WANTED
MOMENT OF
STOWED EQUIPMENT

KNOWN

BL
32.9

EQUIPMENT WEIGHT
= 125 POUNDS

METHOD
ENTER WEIGHT AT
125 POUNDS − MOVE
RIGHT TO LINE
MOVE DOWN READ
MOMENT / 1000 = 52

STA
398.0

STA
420.8

STA
443.5

ARM = 420.8
250

WEIGHT ~ POUNDS

200

150

100

110

MOMENT/1000
AA0595

DATA BASIS: CALCULATED

Figure 6-12. Stowage Compartment Moments

6-23

TM 1-1520-237-10

Section VII CENTER OF GRAVITY
6.21 CENTER OF GRAVITY LIMITS CHART.
The CG limit charts (Figures 6-13 and 6-14) allow the
center of gravity (inches) to be determined when the total
weight and total moment are known.

6-24

Change 5

TM 1-1520-237-10

CENTER OF GRAVITY
WITHOUT EXTERNAL STORES SUPPORT SYSTEM OR
VOLCANO MULTIPLE MINE DELIVERY SYSTEM INSTALLED
11,500 TO 16,500 POUNDS GROSS WEIGHT
CENTER OF GRAVITY LIMITS
EXAMPLE
MAIN ROTOR
CL

16.5

345.8

364.2

DETERMINE IF
LOADING LIMITS
ARE EXCEEDED

580

345.5

KNOWN

560

540

GROSS WEIGHT ~ 1000 POUNDS

ENTER GROSS
WEIGHT AT 15,000
POUNDS. MOVE
RIGHT TOTAL
MOMENT / 1,000
= 5,400 CG
IS WITHIN LIMITS
MOVE DOWN TO
ARM = 360

EXAMPLE

METHOD

FORWAR

D LIMITS

GROSS WEIGHT
= 15,000 POUNDS
MOMENT / 1000
= 5,400

/ 10

15,900

AFT LIMITS

WANTED

520

500
0
TO
TAL
MO
ME
NTS

342.6

/ 10
00

480
0

13,700

366.3

13,400
460
0

13,050

366.3
440
0

12,500
363.2

420
0

TOT

12,000

MO
M

ENT

400

360.8

100

11.5
335

LEGEND

340

345

350

355

360

365

370

ARM ~ INCHES

BEYOND LIMITS

DATA BASIS:

CALCULATED

AA0347_1B
SA

Figure 6-13. Center of Gravity Limits Chart (Sheet 1 of 2)

6-25

TM 1-1520-237-10

CENTER OF GRAVITY
WITHOUT EXTERNAL STORES SUPPORT SYSTEM OR
VOLCANO MULTIPLE MINE DELIVERY SYSTEM INSTALLED
16,000 TO 23,500 POUNDS GROSS WEIGHT
CENTER OF GRAVITY LIMITS
MAXIMUM GROSS WEIGHT
FOR UH−60L EXTERNAL
LOAD MISSION FOR CARGO
HOOK LOADS ABOVE
8,000 LBS UP TO
9,000 LBS.

MAIN ROTOR CL
23.5

348.2

359.2
23,500
840

23
80

22,000

MAXIMUM GROSS WEIGHT
FOR ALL UH−60L AND
SOME UH / EH−60A.
SEE PARAGRAPH 5.14
FOR DETAILS.

20,250
20

AFT LIMITS

FORWARD LIMITS

MAXIMUM GROSS WEIGHT
FOR SOME UH / EH−60A.
SEE PARAGRAPH 5.14
FOR DETAILS.

GROSS WEIGHT ~ 1000 POUNDS

66
TO

64
00

18
620

600

16.5
560

16
335

DATA BASIS: CALCULATED

/ 10

LEGEND
BEYOND LIMITS

580

340

345

350

355

360

365

ARM ~ INCHES

AA8801A
SA

Figure 6-13. Center of Gravity Limits Chart (Sheet 2 of 2)
6-26

370

TM 1-1520-237-10

CENTER OF GRAVITY
WITH EXTERNAL STORES SUPPORT SYSTEM INSTALLED
11,500 TO 16,500 POUNDS GROSS WEIGHT
CENTER OF GRAVITY LIMITS

MAIN ROTOR
CL

16.5

364.2

580

560

/ 10

0
AFT LIMITS

540

520

FORWARD LIMITS

GROSS WEIGHT ~ 1000 POUNDS

500
0
TO
TAL
MO
ME
NTS

/ 10
00

480
0

366.3

0
13,400
460
0

13,050

366.3
440
0

12,500

0
363.2
420
0

TOT

12,000

MO
M

ENT

400
0

360.8

100

11.5
335

LEGEND

340

345

350

355

360

365

370

ARM ~ INCHES

BEYOND LIMITS
AA1254_1B

DATA BASIS: CALCULATED

Figure 6-14. Center of Gravity Limits Chart (Sheet 1 of 3)

6-27

TM 1-1520-237-10

CENTER OF GRAVITY
WITH EXTERNAL STORES SUPPORT SYSTEM OR
VOLCANO MULTIPLE MINE DELIVERY SYSTEM INSTALLED
16,000 TO 22,000 POUNDS GROSS WEIGHT
CENTER OF GRAVITY LIMITS
MAIN ROTOR CL 343.0

360.2

MAXIMUM GROSS WEIGHT
FOR ALL UH−60L AND
SOME UH / EH−60A.
SEE PARAGRAPH 5.14
FOR DETAILS.

343.0

21,500

FORWARD LIMITS

20,250
70

GROSS WEIGHT ~ 1000 POUNDS

AFT LIMITS

MAXIMUM GROSS WEIGHT
FOR SOME UH / EH−60A.
SEE PARAGRAPH 5.14
FOR DETAILS.

620

341.0

600

16,825

/ 10

580
0

16.5
560
0

LEGEND
16

BEYOND LIMITS

DATA BASIS:

335

CALCULATED

340

345

350

355

360

ARM ~ INCHES

Figure 6-14. Center of Gravity Limits Chart (Sheet 2 of 3)

6-28

365

370
AA1254_2C
SA

TM 1-1520-237-10

CENTER OF GRAVITY
WITH EXTERNAL STORES SUPPORT SYSTEM OR
VOLCANO MULTIPLE MINE DELIVERY SYSTEM INSTALLED
21,750 TO 24,500 POUNDS GROSS WEIGHT
CENTER OF GRAVITY LIMITS

MAIN ROTOR
CL
343

345

24.5

GROSS WEIGHTS ABOVE THE MAXIMUM
VALUES SPECIFIED IN PARAGRAPH 5.14
ARE LIMITED TO FERRY MISSIONS
FOR WHICH AN AIRWORTHINESS
RELEASE IS REQUIRED

S
AFT LIMIT

FORWARD LIMITS

GROSS WEIGHT ~ 1000 POUNDS

83
00

78
00

360.2

76
00

347.7
22,000

78
00
21.75
335

340

345

350

355

360

365

370

ARM ~ INCHES

LEGEND
BEYOND LIMITS

DATA BASIS:

CALCULATED

AA1254_3B
SA

Figure 6-14. Center of Gravity Limits Chart (Sheet 3 of 3)

6-29/(6-30 Blank)

TM 1-1520-237-10

CHAPTER 7
PERFORMANCE DATA
Section I INTRODUCTION
NOTE
Chapter 7 contains performance data for aircraft equipped with T700-GE-700 engines.
Performance data for other models are contained in Chapter 7A. Users are authorized
to remove whichever chapter is not applicable to their model aircraft, and are not required to carry both chapters on board.
Tabular performance data is presented in the
checklist (TM 1 -1520-237-CL) and may be
used in lieu of Figures 7-3 and 7-4 to obtain
9Maximum Hover Weight9, 9Torque Required to Hover9 and 9Maximum Torque
Available9.
7.1 PURPOSE.
a. The purpose of this chapter is to provide the best
available performance data. Regular use of this information
will enable you to receive maximum safe utilization of the
helicopter. Although maximum performance is not always
required, regular use of this chapter is recommended for
these reasons:

also be used in flight, to establish unit or area standard
operating procedures, and to inform ground commanders of
performance/risk trade-offs.
7.2 CHAPTER 7 INDEX.
The following index contains a list of the sections, titles,
figure numbers, subjects and page numbers of each performance data chart contained in this chapter.

Section
and
Figure
Number

(2) Situations requiring maximum performance will be
more readily recognized.
(3) Familiarity with the data will allow performance to
be computed more easily and quickly.
(4) Experience will be gained in accurately estimating
the effects of variables for which data are not presented.
b. The information is primarily intended for mission
planning and is most useful when planning operations in
unfamiliar areas or at extreme conditions. The data may

Page

INTRODUCTION ...................

7-1

Temperature Conversion
Chart.........................................

7-4

MAXIMUM TORQUE
AVAILABLE...........................

7-6

Aircraft Torque Factor
(ATF) .......................................

7-7

Maximum Torque Available 30-Minute Limit ......................

7-8

III

HOVER....................................

7-9

7-4

Hover - Clean
Configuration ...........................

7-10

7-5

Hover - High Drag ..................

7-12

CRUISE ...................................

7-13

7-6

Sample Cruise Chart -Clean....

7-15

7-7

Cruise - Pressure Altitude Sea
Level ........................................

7-16

Cruise High Drag - Pressure
Altitude Sea Level...................

7-22

Cruise - Pressure Altitude
2,000 Feet ................................

7-28

II
7-2
7-3

(1) Knowledge of your performance margin will allow
you to make better decisions when unexpected conditions
or alternate missions are encountered.

Title

7-8
7-9

Change 6

7-1

TM 1-1520-237-10

Section
and
Figure
Number
7-10
7-11

7-12
7-13
7-14
7-15
7-16
7-17
7-18
7-19
7-20
7-21
7-22
7-23
7-24
7-25
7-26
7-27
7-28
V

7-2

Title

Page

Cruise High Drag - Pressure
Altitude 2,000 Feet..................

7-34

Cruise - Pressure Altitude
4,000 Feet ................................

7-40

Section
and
Figure
Number
7-29

Page

Optimum Altitude for
Maximum Range .....................

7-133

DRAG ......................................

7-135

7-30

External Load Drag .................

7-136

7-31

Typical High Drag
Configurations..........................

7-137

VII

CLIMB - DESCENT...............

7-138

7-32

Climb/Descent..........................

7-139

7-33

Climb/Descent High Drag ................................

7-140

VIII

FUEL FLOW...........................

7-141

Single/Dual-Engine
Fuel Flow.................................

7-142

AIRSPEED SYSTEM
CHARACTERISTICS .............

7-143

Airspeed Correction
Aircraft Without Wedge
Mounted Pitot-Static Probes....

7-144

Airspeed Correction
Aircraft With Wedge Mounted
Pitot-Static Probes ...................

7-145

Airspeed Correction
Chart - High Drag ...................

7-146

SPECIAL MISSION
PERFORMANCE ....................

7-147

Self Deployment Mission
Profile.......................................

7-148

Assault Mission Profile
(4 - 230 Gallon Tanks)............

7-150

Assault Mission Profile
(2 - 230 Gallon Tanks)............

7-151

Cruise High Drag - Pressure
Altitude 4,000 Feet..................

7-46

Cruise - Pressure Altitude
6,000 Feet ................................

7-52

Cruise High Drag - Pressure
Altitude 6,000 Feet..................

7-58

Cruise - Pressure Altitude
8,000 Feet ................................

7-64

7-34

Cruise High Drag - Pressure
Altitude 8,000 Feet..................

7-70

Cruise - Pressure Altitude
10,000 Feet ..............................

7-76

7-35

Cruise High Drag- Pressure
Altitude 10,000 Feet................

7-81

Cruise - Pressure Altitude
12,000 Feet ..............................

7-86

Cruise High Drag - Pressure
Altitude 12,000 Feet................

7-91

Cruise - Pressure Altitude
14,000 Feet ..............................

7-96

Cruise High Drag - Pressure
Altitude 14,000 Feet................

7-101

Cruise - Pressure Altitude
16,000 Feet ..............................

7-106

Cruise High Drag - Pressure
Altitude 16,000 Feet................

7-111

Cruise - Pressure Altitude
18,000 Feet ..............................

7-115

7-36

7-37
X
7-38
7-39
7-40

7.3 GENERAL.

Cruise High Drag - Pressure
Altitude 18,000 Feet................

7-120

Cruise - Pressure Altitude
20,000 Feet ..............................

7-124

Cruise High Drag - Pressure
Altitude 20,000 Feet................

7-128

OPTIMUM CRUISE ...............

7-132

Change 6

Title

The data presented covers the maximum range of conditions and performance that can reasonably be expected.
In each area of performance, the effects of altitude, temperature, gross weight, and other parameters relating to that
phase of flight are presented. In addition to the presented
data, your judgment and experience will be necessary to

TM 1-1520-237-10

accurately obtain performance under a given set of circumstances. The conditions for the data are listed under the title
of each chart. The effects of different conditions are discussed in the text accompanying each phase of perfor-

mance. Where practical, data are presented at conservative
conditions. However, NO GENERAL CONSERVATISM
HAS BEEN APPLIED. All performance data presented are
within the applicable limits of the helicopter. All flight per-

Change 6

7-2.1/(7-2.2 Blank)

TM 1-1520-237-10

formance data are based on JP-4 fuel. The change in fuel
flow and torque available, when using JP-5 or JP-8 aviation
fuel, or any other approved fuels, is insignificant.
7.4 LIMITS.

CAUTION

7.6 PERFORMANCE DISCREPANCIES.
Regular use of this chapter will allow you to monitor
instrument and other helicopter systems for malfunction, by
comparing actual performance with planned performance.
Knowledge will also be gained concerning the effects of
variables for which data is not provided, thereby increasing
the accuracy of performance predictions.
7.7 PERFORMANCE DATA BASIS - CLEAN.

Exceeding operating limits can cause permanent damage to critical components.
Overlimit operation can decrease performance, cause early failure, or failure on a
subsequent flight.
Applicable limits are shown on the charts. Performance
generally deteriorates rapidly beyond limits. If limits are
exceeded, minimize the amount and time. Enter the maximum value and time above limits on DA Form 2408-13-1,
so proper maintenance action can be taken.
7.5 USE OF CHARTS.
7.5.1 Dashed Line Data. Weights above 22,000 pounds
are limited to ferry missions for which an Airworthiness
Release is required. On some charts dashed line data are
shown for gross weights greater than 22,000 pounds.
7.5.2 Data Basis. The type of data used is indicated at
the bottom of each performance chart under DATA BASIS.
The data provided generally is based on one of three categories:
a. Flight test data. Data obtained by flight test of the
helicopter by experienced flight test personnel at precise
conditions using sensitive calibrated instruments.
b. Calculated data. Data based on tests, but not on flight
test of the complete helicopter.
c. Estimated data. Data based on estimates using aerodynamic theory or other means but not verified by flight
test.

The data presented in the performance charts are primarily derived for a clean UH-60A aircraft and are based on
U. S. Army test data. The clean configuration assumes all
doors and windows are closed and includes the following
external configuration:
a. Fixed provisions for the External Stores Support System (ESSS).
b. Main and tail rotor deice system.
c. Mounting brackets for IR jammer and chaff dispenser.
d. The Hover Infrared Suppressor System (HIRSS) with
baffles installed.
e. Includes wire strike protection system.
NOTE
Aircraft which have an external configuration which differs from the clean configuration may be corrected for drag differences
on cruise performance as discussed in Section VI DRAG.
7.8 PERFORMANCE DATA BASIS - HIGH DRAG.
The data presented in the high drag performance charts
are primarily derived for the UH-60A with the ESSS system installed and the 230-gallon tanks mounted on the outboard pylons, and are based on U. S. Army test data. The
high drag configuration assumes all doors and windows are
closed and includes the following external configuration:
a. External stores support system installed.

7.5.3 Specific Conditions. The data presented is accurate only for specific conditions listed under the title of
each chart. Variables for which data is not presented, but
which may affect that phase of performance, are discussed
in the text. Where data is available or reasonable estimates
can be made, the amount that each variable affects performance will be given.

b. Two 230-gallon tanks mounted on the outboard pylons.
c. Inboard vertical pylons empty.
d. IR jammer and chaff dispenser installed.

Change 3

7-3

TM 1-1520-237-10

TEMPERATURE CONVERSION
EXAMPLE
WANTED:
FREE AIR TEMPERATURE IN DEGREES CELSIUS

KNOWN:
FREE AIR TEMPERATURE = 32oF

METHOD:
ENTER FREE AIR TEMPERATURE HERE
MOVE RIGHT TO DIAGONAL LINE
MOVE DOWN TO DEGREES CELSIUS SCALE
READ FREE AIR TEMPERATURE = 0oC

140

120

100

FAT ~ oF

−20

−40

−60

−80
−60

−50

−40

−30

−20

−10

FAT ~ oC
AA0674
SA

Figure 7-1. Temperature Conversion Chart

7-4

TM 1-1520-237-10

e. Hover Infrared Suppressor System (HIRSS) with
baffles are installed.
f. Main and tail rotor deice and wire strike protection
systems are installed.
NOTE
Aircraft with an external configuration that
differs from the high drag configuration
baseline may be corrected for differences in
cruise performance as discussed in Section
VI DRAG.
g. VOL Use the high drag configuration hover charts
to determine hover performance with the volcano system

installed. Use the high drag cruise charts and the volcano
drag correction factor to determine cruise performance with
volcano installed. The volcano drag correction factor is
based on flight test data obtained with the complete volcano
system installed, to include all of the canisters and mines.
The drag correction factor may be used to provide a conservative estimate of cruise performance for volcano configurations which do not include all of the canisters and
mines.
7.9 FREE AIR TEMPERATURES.
A temperature conversion chart (Figure 7-1) is included
for the purpose of converting Fahrenheit temperature to
Celsius.

Change 3

7-5

TM 1-1520-237-10

Section II MAXIMUM TORQUE AVAILABLE
7.10 TORQUE FACTOR METHOD.

7.11 MAXIMUM TORQUE AVAILABLE CHART.

The torque factor method provides an accurate indication of available power by incorporating ambient temperature effects on degraded engine performance. This section
presents the procedure to determine the maximum dual- or
single-engine torque available for the T700-GE-700 engine
as installed in each individual aircraft. Specification power
is defined for a newly delivered low time engine. The aircraft HIT log forms for each engine, provide the engine and
aircraft torque factors which are obtained from the maximum power check and recorded to be used in calculating
maximum torque available.

This chart (Figure 7-3) presents the maximum specification torque available at zero airspeed and 100% RPM R for
the operational range of pressure altitude and FAT. The
single- and dual-engine transmission limits for continuous
operation are shown and should not be exceeded. The engine torque available data above the single-engine transmission limit is presented as dashed lines and is required
for determining torque available when TR values are below
1.0. When the TR equals 1.0, the maximum torque available may be read from the horizontal specification torque
available per engine scale. When the TR value is less than
1.0, the maximum torque available is determined by multiplying the TR by the specification torque available. The
lower portion of Figure 7-3 presents TR correction lines
which may be used in place of multiplication to read torque
available per engine directly from the vertical scale.

7.10.1 Torque Factor Terms. The following terms are
used when determining the maximum torque available for
an individual aircraft:
a. Torque Ratio (TR). The ratio of torque available to
specification torque at the desired ambient temperature.
b. Engine Torque Factor (ETF). The ratio of an individual engine torque available to specification torque at reference temperature of 35°C. The ETF is allowed to range
from 0.85 to 1.0.
c. Aircraft Torque Factor (ATF). The ratio of an individual aircraft’s power available to specification power at a
reference temperature of 35°C. The ATF is the average of
the ETF’s of both engines and its value is allowed to range
from 0.9 to 1.0.
7.10.2 Torque Factor Procedure. The use of the ATF
or ETF to obtain the TR from Figure 7-2 for ambient temperatures between -15°C and 35°C is shown by the example. The ATF and ETF values for an individual aircraft
are found on the engine HIT Log. The TR always equals
1.0 for ambient temperatures of -15°C and below, and the
TR equals the ATF or ETF for temperatures of 35°C and
above. For these cases, and for an ATF or ETF value of 1.0,
Figure 7-2 need not be used.

7-6

7.12 ENGINE BLEED AIR.
With engine bleed air turned on, the maximum available
torque is reduced as follows:
a. Engine Anti-Ice On: Reduce torque determined from
Figure 7-3 by a constant 16% TRQ. Example: (90% TRQ16% TRQ) = 74% TRQ.
b. Cockpit Heater On: Reduce torque available by 4%
TRQ.
c. Both On: Reduce torque available by 20% TRQ.
7.13 INFRARED SUPPRESSOR SYSTEM.
When the hover IR suppressor system is installed and
operating in the benign mode exhaust (baffles removed) the
maximum torque available is increased about 1% TRQ.
When an IR suppressor system is not installed, maximum
torque available is also increased about 1%.

TM 1-1520-237-10

TORQUE FACTOR
TORQUE FACTOR ~ ATF OR ETF
.84

.86

.90

.88

.92

.94

.96

.98

1.0

FOR FAT’S
OF 35oC
AND ABOVE
TR = ATF

FREE AIR TEMPERATURE ~ o C

−5

−10

−15

−20
.84

FOR FAT’S
OF −15OC
AND BELOW
TR = 1.0
.85

.86

.87

.88

.89

.90

.91

.92

.93

.94

.95

.96

.97

.98

.99

1.00

TORQUE RATIO = TR

EXAMPLE
WANTED:
TORQUE RATIO AND MAXIMUM TORQUE AVAILABLE

TO CALCULATE MAXIMUM TORQUE AVAILABLE:

KNOWN:

4. ENTER MAXIMUM TORQUE AVAILABLE CHART AT
KNOWN FAT (FIGURE 7−3)
5. MOVE RIGHT TO KNOWN PRESSURE ALTITUDE
6. MOVE DOWN, READ SPECIFICATION TORQUE = 97.2%

ATF = .95
PRESSURE ALTITUDE = 6000 FT
FAT = 6oC

TO OBTAIN VALUE FROM CHART:

METHOD:
7. MOVE DOWN TO TORQUE RATIO OBTAINED FROM FIGURE 7−2
TO OBTAIN TORQUE RATIO:
8. MOVE LEFT, READ MAXIMUM TORQUE AVAILABLE = 93.0%
1. ENTER TORQUE FACTOR CHART AT KNOWN FAT
2. MOVE RIGHT TO THE ATF VALUE
3. MOVE DOWN, READ TORQUE RATIO = .972

DATA BASIS:

CALCULATED

AA0675A
SA

Figure 7-2. Aircraft Torque Factor (ATF)

7-7

TM 1-1520-237-10

MAXIMUM TORQUE AVAILABLE
100% RPM R
30 MIN LIMIT
ZERO AIRSPEED

HIRSS (BAFFLES INSTALLED)
BLEED AIR OFF

FREE AIR TEMPERATURE ~ oC

40
30

ENGINE HIGH AMBIENT
TEMPERATURE LIMIT
10
12
14

20
10
4

PRESSURE
ALTITUDE
~ 1000 FT

TRANSMISSION LIMITS
2~ENGINE 1~ENGINE

2
4
6
8

16
18

0
−10
−20
6
−30
−40
−50
50

100

110

120

SPECIFICATION TORQUE AVAILABLE PER ENGINE %
110

1.00

TRANSMISSION LIMIT ~ 1 ENGINE

0.96

0.92
0.88

TRANSMISSION LIMIT ~ 2 ENGINE

0.84

100

TORQUE AVAILABLE ~ %

TORQUE RATIO

60
SPECIFIC TORQUE
X TORQUE RATIO
= TORQUE AVAILABLE
50

DATA BASIS:

AA0381B

FLIGHT TEST

Figure 7-3. Maximum Torque Available - 30-Minute Limit
7-8

TM 1-1520-237-10

Section III HOVER
7.14 HOVER CHART.
NOTE
VOL For performance calculations with
volcano system installed, use the applicable
high drag performance charts.

a. The primary use of the chart (Figures 7-4 through
7-5) is illustrated by part A of the example. To determine
the torque required to hover, it is necessary to know pressure altitude, free air temperature, gross weight, and desired
wheel height. Enter the upper right grid at the known free
air temperature, move right to the pressure altitude, move
down to gross weight. For OGE hover, move left to the
torque per engine scale and read torque required. For IGE
hover, move left to desired wheel height, deflect down and
read torque required for dual-engine or single-engine operation. The IGE wheel height lines represent a compromise for all possible gross weights and altitude conditions.
A small torque error up to 63% torque may occur at extreme temperature and high altitude. This error is more evident at lower wheel heights.
b. In addition to the primary use, the hover chart (Figure
7-4) may be used to predict maximum hover height. To
determine maximum hover height, it is necessary to know
pressure altitude, free air temperature, gross weight, and
maximum torque available. Enter the known free air temperature move right to the pressure altitude, move down to

gross weight, move left to intersection with maximum
torque available and read wheel height. This wheel height
is the maximum hover height.
c. The hover chart may also be used to determine maximum gross weight for hover at a given wheel height, pressure altitude, and temperature as illustrated in method B of
the example (Figure 7-4). Enter at known free air temperature, move right to the pressure altitude, then move down
and establish a vertical line on the lower grid. Now enter
lower left grid at maximum torque available. Move up to
wheel height, then move right to intersect vertical line from
pressure altitude/FAT intersection. Interpolate from gross
weight lines to read maximum gross weight at which the
helicopter will hover.
7.15 EFFECTS OF BLADE EROSION KIT.
With the blade erosion kit installed, it will be necessary
to make the following corrections. Multiply the torque required to hover determined from the charts by 1.02. (Example: If indicated torque is 90%, multiply 90 x 1.02 =
91.8% actual torque required.) Multiply the maximum gross
weight to hover obtained from the charts by 0.98. (Example: If gross weight is 22,000 lb, multiply by 0.98 =
21,560 lb actual gross weight to hover.) When determining
maximum hover wheel height, enter the chart at 1.02 x
gross weight. (Example: If gross weight is 20,000 lb, multiply 20,000 x 1.02 = 20,400 lb).

7-9

TM 1-1520-237-10

EXAMPLE A
WANTED:
TORQUE REQUIRED TO HOVER OGE AND AT A 10-FOOT WHEEL HEIGHT
KNOWN:
FAT = 30°C
PRESSURE ALTITUDE = 2,000 FEET
GROSS WEIGHT = 19,500 POUNDS
METHOD:
ENTER HOVER CHART AT KNOWN FAT. MOVE RIGHT TO PRESSURE ALTITUDE, MOVE DOWN
THROUGH GROSS WEIGHT LINES TO DESIRED GROSS WEIGHT. MOVE LEFT TO INDICATE
TORQUE/ENGINE % (OGE) SCALE AND READ OGE HOVER TORQUE (94%). MOVE DOWN
FROM INTERSECTION OF 10-FOOT HOVER LINE AND HORIZONTAL LINE TO READ TORQUE
REQUIRED TO HOVER 10 FEET (80%).

EXAMPLE B
WANTED:
MAXIMUM GROSS WEIGHT TO HOVER OGE
KNOWN:
ATF = 1.0
FAT = 15°C
PRESSURE ALTITUDE = 8,000 FEET
MAXIMUM TORQUE AVAILABLE = 96%
METHOD:
ENTER INDICATED TORQUE/ENGINE (IGE) SCALE AT MAXIMUM TORQUE AVAILABLE (96%),
MOVE UP TO OGE LINE. ENTER CHART AT KNOWN FAT (15°C). MOVE RIGHT TO PRESSURE
ALTITUDE LINE (8,000 FT). MOVE DOWN FROM PRESSURE ALTITUDE LINE AND MOVE
RIGHT FROM OGE LINE. WHERE LINES INTERSECT, READ MAXIMUM GROSS WEIGHT TO
HOVER OGE.
Figure 7-4. Hover - Clean Configuration (Sheet 1 of 2)

7-10

TM 1-1520-237-10

HOVER

HOVER
CLEAN
T700(2)

CLEAN CONFIGURATION 100% RPM R
ZERO WIND

PRESSURE ALTITUDE ~ 1000 FT

FREE AIR TEMP. ~ OC

−2

NOTE
FOR LOW WIND CONDITIONS
AIRCRAFT SHOULD BE HEADED
INTO WIND. 3−5 KT CROSSWIND
OR TAILWIND MAY INCREASE
TORQUE REQUIRED BY UP TO
4% OVER ZERO WIND VALUES

16
18

A
20
B
0

−20
−40
−60

WHEEL
HEIGHT ~FT

40
100

DUAL ENGINE TRANS. LIMIT

15
OGE

95
SINGLE ENGINE
TRANS. LIMT

14
85
80

DUAL ENGINE TRANS. LIMIT

TORQUE PER ENGINE ~ % (OGE)

75
70
65

60
55
B

GROSS
WEIGHT
~ 1000 LB

50
40

100

TORQUE PER ENGINE ~ % (IGE)
80

DATA BASIS:

100

120

140

SINGLE ENGINE TORQUE ~ %
AA2143C

FLIGHT TEST

Figure 7-4. Hover - Clean Configuration (Sheet 2 of 2)

Change 2

7-11

TM 1-1520-237-10

HOVER

HOVER
ESSS
T700 (2)

HIGH DRAG CONFIGURATION 100% RPM R
ZERO WIND

PRESSURE ALTITUDE ~ 1000 FT

FREE AIR TEMP. ~ OC

−2

NOTE
FOR LOW WIND CONDITIONS
AIRCRAFT SHOULD BE HEADED
INTO WIND. 3−5 KT CROSSWIND
OR TAILWIND MAY INCREASE
TORQUE REQUIRED BY UP TO
4% OVER ZERO WIND VALUES

20
20

0
−20
−40
−60

WHEEL
HEIGHT ~FT

24.5 24

20
15
40
DUAL ENGINE TRANS. LIMIT

100

OGE
SINGLE ENGINE
TRANS. LIMIT

DUAL ENGINE TRANS. LIMIT

TORQUE PER ENGINE ~ % (OGE)

85
80
75

70
65

GROSS
WEIGHT
~ 1000 LB

60
55
40

100

TORQUE PER ENGINE ~ % (IGE)
80

100

DATA BASIS:

120

140

SINGLE ENGINE TORQUE ~ %

FLIGHT TEST

AA2144C
SA

Figure 7-5. Hover - High Drag

7-12

Change 2

TM 1-1520-237-10

Section IV CRUISE
7.16 DESCRIPTION.
The cruise charts (Figures 7-6 through 7-28) present
torque required and total fuel flow as a function of airspeed,
altitude, temperature, and gross weight at 100% rotor speed.
Scales for both true airspeed and indicated airspeed are
presented. The baseline aircraft configuration for these
charts was the 9clean and high drag9 configuration as defined in Section I. Each cruise chart also presents the
change in torque ( TRQ) required for 10 sq. ft. of additional flat plate drag with a dashed line on a separate scale.
This line is utilized to correct torque required for external
loads as discussed in Section VI DRAG. Maximum level
flight airspeed (Vh) is obtained at the intersection of gross
weight arc and torque available - 30 minutes or the transmission torque limit, whichever is lower. Airspeeds that
will produce maximum range, maximum endurance, and
maximum rate of climb are also shown. Cruise charts are
provided from sea level to 20,000 feet pressure altitude in
units of 2,000 feet. Each figure number represents a different altitude. The charts provide cruise data for free air temperatures from -50° to +60°C, in units of 10°. Charts with
FAT’s that exceed the engine ambient temperature limits
by more than 10°C are deleted.
7.17 USE OF CHARTS.
The primary uses of the charts are illustrated by the
examples of Figure 7-6. To use the charts, it is usually
necessary to know the planned pressure altitude, estimated
free air temperature, planned cruise speed, TAS, and gross
weight. First, select the proper chart on the basis of pressure altitude and FAT. Enter the chart at the cruise airspeed, IAS, move horizontal and read TAS, move horizontal to the gross weight, move down and read torque
required, and then move up and read associated fuel flow.
Maximum performance conditions are determined by entering the chart where the maximum range line or the maximum endurance and rate of climb line intersects the gross
weight line; then read airspeed, fuel flow, and torque required. Normally, sufficient accuracy can be obtained by
selecting the chart nearest the planned cruising altitude and
FAT or, more conservatively, by selecting the chart with
the next higher altitude and FAT. If greater accuracy is
required, interpolation between altitudes and/or temperatures is permissible. To be conservative, use the gross
weight at the beginning of the cruise flight. For greater
accuracy on long flights, however, it is preferable to determine cruise information for several flight segments to allow
for the decreasing gross weight.

a. Airspeed. True and indicated airspeeds are presented
at opposite sides of each chart. On any chart, indicated
airspeed can be directly converted to true airspeed (or vice
versa) by reading directly across the chart without regard
for the other chart information. The level flight airspeed
calibration for aircraft with wedge mounted pitot static
probes (hard points only) was used to convert indicated to
true airspeed.
b. Torque. Since pressure altitude and temperature are
fixed for each chart, torque required varies according to
gross weight and airspeed. The torque and torque limits
shown on these charts are for dual-engine operation. The
maximum torque available is presented on each chart as
either the transmission torque limit or torque available - 30
minute for both ATF-1.0 and 0.9 values. The maximum
torque available for aircraft with an ATF value between
these shall be interpolated. The continuous torque available
values shown represent the minimum torque available for
ATF’s of 0.95 or greater. For ATF’s less than 0.95 maximum continuous torque available may be slightly reduced.
Higher torque than that represented by these lines may be
used if it is available without exceeding the limitations presented in Chapter 5. An increase or decrease in torque required because of a drag area change is calculated by adding or subtracting the change in torque from the torque on
the curve, and then reading the new fuel flow total.
c. Fuel Flow. Fuel flow scales are provided opposite the
torque scales. On any chart, torque may be converted directly to fuel flow without regard to other chart information. Data shown in this section is for two-engine operation.
For one-engine fuel flow, refer to Section VIII FUEL
FLOW.
(1) With bleed-air extracted, fuel flow increases:
(a) Engine anti-ice on -About 60 lbs/hr Example:
(760 lbs/hr + 60 lbs/hr = 820 lbs/hr.)
(b) Heater on - About 20 lbs/hr.
(c) Both on - About 80 lbs/hr.
(2) When the cruise IR suppressors are removed or
hover IR suppressor system is installed and operating in the
benign mode (exhaust baffles removed), the dual-engine
fuel flow will decrease about 16 lbs/hr.
d. Maximum Range. The maximum range lines (MAX
RANGE) indicate the combinations of gross weight and

7-13

TM 1-1520-237-10

airspeed that will produce the greatest flight range per
pound of fuel under zero wind conditions. When maximum
range airspeed line is above the maximum torque available,
the resulting maximum airspeed should be used for maximum range. A method of estimating maximum range speed
in winds is to increase IAS by 2.5 knots per each 10 knots
of effective headwind (which reduces flight time and minimizes loss in range) and decrease IAS by 2.5 knots per 10
knots of effective tailwind for economy.
e. Maximum Endurance and Rate of Climb. The maximum endurance and rate of climb lines (MAX END and
R/C) indicate the combinations of gross weight and airspeed that will produce the maximum endurance and the
maximum rate of climb. The torque required for level flight
at this condition is a minimum, providing a minimum fuel
flow (maximum endurance) and a maximum torque change
available for climb (maximum rate of climb).

ally result in cruise at best range airspeed for the higher
drag configuration. To determine the approximate airspeed
for maximum range for alternative or external load configurations, reduce the value from the cruise chart by 6 knots
for each 10 square foot increase in drag area,
F. For
example, if both cabin doors are open the F increases 6
ft2 and the maximum range airspeed would be reduced by
approximately 4 knots (6 Kts/10 ft236 ft2 = 3.6 Kts).
g. Additional Uses. The low speed end of the cruise
chart (below 40 knots) is shown primarily to familiarize
you with the low speed power requirements of the helicopter. It shows the power margin available for climb or acceleration during maneuvers, such as NOE flight. At zero airspeed, the torque represents the torque required to hover
out of ground effect. In general, mission planning for low
speed flight should be based on hover out of ground effect.
7.18 SINGLE-ENGINE.

f. Change in Frontal Area. Since the cruise information
is given for the 9clean configuration,9 adjustments to torque
should be made when operating with external sling loads or
aircraft external configuration changes. To determine the
change in torque, first obtain the appropriate multiplying
factor from the drag load chart (Figure 7-30), then enter the
cruise chart at the planned cruise speed TAS, move right to
TRQ line, and move up and read
TRQ.
the broken
Multiply
TRQ by the multiplying factor to obtain change
in torque, then add or subtract change in torque from torque
required for the primary mission configuration. Enter the
cruise chart at resulting torque required, move up, and read
fuel flow. If the resulting torque required exceeds the governing torque limit, the torque required must be reduced to
the limit. The resulting reduction in airspeed may be found
by subtracting the change in torque from the limit torque;
then enter the cruise chart at the reduced torque, and move
up to the gross weight. Move left or right to read TAS or
IAS. The engine torque setting for maximum range obtained from the clean configuration cruise chart will gener-

7-14

a. The minimum or maximum single-engine speeds can
be determined by using a combination of the 700 torque
available and cruise charts. To calculate single-engine
speeds, first determine the torque available from Section II
at the TGT limit desired and divide by 2. (Example: 90%
TRQ42 = 45% TRQ.)
b. Select the appropriate cruise chart for the desired
flight condition and enter the torque scale with the torque
value derived above. Move up to the intersection of torque
available and the mission gross weight arc, and read across
for minimum single-engine airspeed. Move up to the second intersection of torque and weight, and read across to
determine the maximum single-engine speed. If no intersections occur, there is no single-engine level flight capability for the conditions. Single-engine fuel flow at the desired 10 minute, 30 minute, continuous conditions may be
obtained by doubling the torque required from the cruise
chart and referring to Figure 7–34.

TM 1-1520-237-10

CRUISE EXAMPLE
CLEAN CONFIGURATION
100% RPM R
FAT: 30 OC ALT: 6,000 FT
TOTAL FUEL FLOW 100 LB/HR

EXAMPLE
6

WANTED:

13
170

180

~ CONTINUOUS

170

KNOWN:
160

FAT = + 30oC
PRESSURE ALTITUDE = 6000 FT
GW = 17000 LBS
ATF = 0.95

ATF = 0.9

TRQ ~ % OR DRAG
AREA OF 10 SQ FT

ATF = 1.0

A. CRUISE CONDITIONS FOR MAXIMUM RANGE
B. CONDITIONS FOR MAXIMUM ENDURANCE
C. MAXIMUM AIRSPEED IN LEVEL FLIGHT
D. DETERMINE TORQUE AND FUEL FLOW
REQUIRED TO CRUISE AT THE CONDITIONS
OF EXAMPLE A WITH CABIN DOORS OPEN

160

150

140
C

150
130
140
MAX
RANGE

D. ENTER TRQ% PER 10 SQ FT SCALE AT 135 KTAS
MOVE UP READ TRQ = 8.0%
TURN TO DRAG TABLE IN SECTION VII
NOTE CABIN DOORS OPEN = 6.0 SQ FT F
AND HAS A DRAG MULTIPLYING FACTOR VALUE
OF 0.60, CALCULATE TOTAL TORQUE REQUIRED:

120

110

100

MAX END
AND R / C

TORQUE AVAILABLE ~ 30 MINUTES

C. AT INTERSECTION OF 30−MINUTE TORQUE
AVAILABLE AS INTERPOLATED FOR THE ATF
VALUE AT THE KNOWN GROSS WEIGHT:
MOVE LEFT, READ MAXIMUM TAS = 153 KTS
MOVE RIGHT, READ MAXIMUM IAS = 135 KTS
MOVE DOWN, READ MAXIMUM TORQUE = 82% TRQ
MOVE UP, READ TOTAL FUEL FLOW = 1125 LBS / HR

130

TRUE AIRSPEED ~ KTS

B. AT INTERSECTION OF MAX END. / AND R / C
LINE AND KNOWN VALUE OF GROSS WEIGHT:
MOVE LEFT, READ TAS = 82 KTS
MOVE RIGHT, READ IAS = 67 KTS
MOVE DOWN, READ TORQUE = 41% TRQ
MOVE UP, READ TOTAL FUEL FLOW = 700 LBS / HR

120
A
110

100

INDICATED AIRSPEED ~ KTS

A. TURN TO CRUISE CHARTS NEAREST KNOWN
FLIGHT CONDITIONS, AT INTERSECTION
OF MAX RANGE LINE AND KNOWN VALUE OF
GROSS WEIGHT:
MOVE LEFT, READ TAS = 135 KTS
MOVE RIGHT, READ IAS = 119 KTS
MOVE DOWN, READ TORQUE = 62% TRQ
MOVE UP, READ TOTAL FUEL FLOW = 900 LBS / HR

TRANSMISSION TORQUE LIMIT

METHOD:

70
50
60
40
50

62% + (0.6 X 8.0%) = 66.8% TRQ

READ FUEL FLOW AT−TOTAL TORQUE = 950 LBS / HR

40
12

22
20

GW ~
1000 LB

10
10
0

0
20

TORQUE PER ENGINE ~ %

100
AA0682
SA

Figure 7-6. Sample Cruise Chart - Clean

7-15

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
140

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

140

150

130

130
140

120
MAX
RANGE

120

MAX
RANGE

130

110
110
120
100

100

80
MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

100

110

30
30

30
GW ~
1000 LB

20
12

GW ~
1000 LB

20
12

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-7. Cruise - Pressure Altitude Sea Level (Sheet 1 of 6)
7-16

AA0414
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140
140

150
130

130

140
MAX
RANGE

120

110

120

110

100
100

90
90
MAX END
AND R / C
TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

100

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

30
30

30
20

GW ~
1000 LB

20
12

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0413
SA

Figure 7-7. Cruise - Pressure Altitude Sea Level (Sheet 2 of 6)
7-17

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

30
150

150

140

140
140

130
MAX
RANGE

130

120

130

MAX
RANGE

120
120
110
110

100

100
90

MAX END
AND R / C

80
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

30
30

30
20
GW ~
1000 LB

GW ~
1000 LB

22
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-7. Cruise - Pressure Altitude Sea Level (Sheet 3 of 6)
7-18

AA0412
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

120

110

100

TRANSMISSION TORQUE LIMIT

100

MAX END
AND R / C

100

80
MAX END
AND R / C
70

40
40

20
20

GW ~
1000 LB

0
20

TORQUE PER ENGINE ~ %

100

TRUE AIRSPEED ~ KTS

110

130

TRANSMISSION TORQUE LIMIT

120

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

~CONTINUOUS

MAX
RANGE

130

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0415
SA

Figure 7-7. Cruise - Pressure Altitude Sea Level (Sheet 4 of 6)
7-19

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

ATF=1.0

40OC
ATF=0.9

30OC

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

140
MAX
RANGE

130

MAX END
AND R / C

110

100

80
MAX END
AND R / C
70

110

100

50
12

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

60
60

TRUE AIRSPEED ~ KTS

120
TORQUE AVAILABLE ~ 30 MINUTES

100

120

~CONTINUOUS

110

TORQUE AVAILABLE ~ 30 MINUTES, ATF=0.9

130

120

TRUE AIRSPEED ~ KTS

MAX
RANGE

40
12

GW ~
1000 LB

40
30

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-7. Cruise - Pressure Altitude Sea Level (Sheet 5 of 6)
7-20

AA0416
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
50OC

60OC

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

180

170

150

140

150

140

MAX
RANGE

ATF=1.0

160

ATF=0.9

~CONTINUOUS

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF=1.0

~CONTINUOUS

170

ATF=0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170
160
150

MAX
RANGE

130

140

130

120
120

120

MAX END
AND R / C

80
MAX END
AND R / C

110
TRANSMISSION TORQUE LIMIT

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MINUTES

110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

110

100

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

22
40

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0417
SA

Figure 7-7. Cruise - Pressure Altitude Sea Level (Sheet 6 of 6)
7-21

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

PRESS ALT: 0 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150

140

140
130

120

130
120
MAX
RANGE

MAX
RANGE

120

110

100
100

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

80
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

110

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

50
40

14 16 18

24.5

12 14 16 18

GW ~
1000 LB

24.5

GW ~
1000 LB

30
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-8. Cruise High Drag - Pressure Altitude Sea Level (Sheet 1 of 6)
7-22

AA0534
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

PRESS ALT: 0 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150
140

140

140
130

130

130
120

120

MAX
RANGE

120

110

100

100
90

80
MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

80
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

MAX
RANGE

50
40
12 14 16 18

24.5

GW ~
1000 LB

14 16 18

24.5
40

GW ~
1000 LB

20
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0535
SA

Figure 7-8. Cruise High Drag - Pressure Altitude Sea Level (Sheet 2 of 6)
7-23

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

PRESS ALT: 0 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

140

130

140
130
MAX
RANGE

110

100

TRANSMISSION TORQUE LIMIT

110

80
MAX END
AND R / C
70

120

110

100

80
MAX END
AND R / C

50
12

14 16 18

24.5

GW ~
1000 LB

14 16

24.5

GW ~
1000 LB

40
40

TRUE AIRSPEED ~ KTS

120

TRANSMISSION TORQUE LIMIT

MAX
RANGE

120

TRUE AIRSPEED ~ KTS

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-8. Cruise High Drag - Pressure Altitude Sea Level (Sheet 3 of 6)
7-24

AA0536
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

PRESS ALT: 0 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

180

170

150

160

140

150

150
140

140

130
130
MAX
RANGE

130

MAX
RANGE

120

TORQUE
AVAILABLE
~ CONTINUOUS

100

80
80

MAX END
AND R / C

50
12

14 16

24.5

GW ~
1000 LB

MAX END
AND R / C

50
12

14 16

24.5

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

100

TRANSMISSION TORQUE LIMIT

100

110

TORQUE
AVAILABLE
~ CONTINUOUS

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0537
SA

Figure 7-8. Cruise High Drag - Pressure Altitude Sea Level (Sheet 4 of 6)
7-25

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

PRESS ALT: 0 FT
30OC

40OC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160

~ CONTINUOUS

140

130

MAX
RANGE

110

100
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

MAX END
AND R / C

150
140

MAX
RANGE

120

130

110

120

100

110

80
MAX END
AND R / C

50
12

24.5

GW ~
1000 LB

23
40

GW ~
1000 LB
20

100

24.5
40

170
160

TORQUE AVAILABLE ~ 30 MINUTES

150

~ CONTINUOUS

160

180

TRUE AIRSPEED ~ KTS

170

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRANSMISSION TORQUE LIMIT

180

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-8. Cruise High Drag - Pressure Altitude Sea Level (Sheet 5 of 6)
7-26

AA0538
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T700 (2)

PRESS ALT: 0 FT
50oC

60oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

CONTINUOUS

170
160
150

180

150

140

130

ATF= 1.0

160

CONTINUOUS

ATF= 1.0

ATF= 0.9

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170
160
150

140

140
MAX
RANGE

130

MAX
RANGE

130

120

120
100

MAX END
AND R/C

TRANSMISSION TORQUE LIMIT

110
TORQUE AVAILABLE ~ 30 MIN

100

TRANSMISSION TORQUE LIMIT

110

100

TRUE AIRSPEED ~ KTS

110

TORQUE AVAILABLE ~ 30 MIN

TRUE AIRSPEED ~ KTS

120

60
40

50
12

24.5

GW ~
1000 LB

24.5
40

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0539
SA

Figure 7-8. Cruise High Drag - Pressure Altitude Sea Level (Sheet 6 of 6)
7-27

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 2000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140
140

150
130

130

140
120

MAX
RANGE

110

120

100

110

100
100

80
MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

90
90
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

50
40

40
12

16 18

GW ~
1000 LB

40
30

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-9. Cruise - Pressure Altitude 2,000 Feet (Sheet 1 of 6)
7-28

AA0449
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 2000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

160
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150

140

140
140

130
MAX
RANGE

120

130

MAX
RANGE

130

120
120
110
110

100

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

100
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

50
50

50
40

16 18

GW ~
1000 LB

16 18

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0448
SA

Figure 7-9. Cruise - Pressure Altitude 2,000 Feet (Sheet 2 of 6)
7-29

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 2000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

MAX
RANGE

130

120

100

120

110

100

MAX END
AND R / C

100

90
90
80

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

120

MAX END
AND R / C

60
50
50
40
40

16 18

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

20
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-9. Cruise - Pressure Altitude 2,000 Feet (Sheet 3 of 6)
7-30

AA0447
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 2000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

140
MAX
RANGE

130

80
MAX END
AND R / C
70

GW ~
1000 LB

GW ~
1000 LB
20

110

100

TRUE AIRSPEED ~ KTS

MAX END
AND R / C

100

120

70
TRANSMISSION TORQUE LIMIT

~ CONTINUOUS

100

110

TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ CONTINUOUS

110

TORQUE AVAILABLE ~ 30 MINUTES ATF= 0.9

120

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

40
30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0450
SA

Figure 7-9. Cruise - Pressure Altitude 2,000 Feet (Sheet 4 of 6)
7-31

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 2000 FT

30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

180
170

150

160

140

150

90
MAX END
AND R / C
80

TES

80
MAX END
AND R / C
70

50
12

110

100

GW ~
1000 LB

GW ~
1000 LB
30

120

50
12

130

TRANSMISSION TORQUE LIMIT

100

TRANSMISSION TORQUE LIMIT

110

120

110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

120

ATF = 1.0

ATF = 0.9

~ CONTINUOUS

130

140

MAX
RANGE

130

TORQUE AVAILABLE ~ 30 MINU

MAX
RANGE

~ CONTINUOUS

140

TRUE AIRSPEED ~ KTS

170

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 0.9

170

180

20
10
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %
AA0451
SA

Figure 7-9. Cruise - Pressure Altitude 2,000 Feet (Sheet 5 of 6)
7-32

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 2000 FT
50oC

60oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160

180
170

150

160

160
140

150

150
130

110

100

TORQUE AVAILABLE ~ 30 MINUTES

100

90
MAX END
AND R / C
80

130

120

80
MAX END
AND R / C

110

100

TRUE AIRSPEED ~ KTS

120

140

TRANSMISSION TORQUE LIMIT

120

TRANSMISSION TORQUE LIMIT

130

MAX
RANGE

TORQUE AVAILABLE ~ 30 MINUTES

MAX
RANGE

140

TRUE AIRSPEED ~ KTS

ATF= 1.0

ATF= 0.9

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 1.0

ATF= 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

70
50

60
40

50
12

22
40

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0452
SA

Figure 7-9. Cruise - Pressure Altitude 2,000 Feet (Sheet 6 of 6)
7-33

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

PRESS ALT: 2000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150
140

140

140
130

130
130
120
120
120
MAX
RANGE

110

MAX
RANGE

110

100

80
MAX END
AND R/C

TRANSMISSION TORQUE LIMIT

100

80
MAX END
AND R/C

TRUE AIRSPEED ~ KTS

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

50
50

50
40

16 18

24.5

GW ~
1000 LB

24.5

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-10. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 1 of 6)
7-34

AA0540
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

PRESS ALT: 2000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

140
130

120

130

120

100

80
MAX END
AND R / C
70

14 16 18 20

24.5

110

100

80
MAX END
AND R / C

50
12

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

120

MAX
RANGE

14 16 18 20

TRUE AIRSPEED ~ KTS

130

24.5

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0541
SA

Figure 7-10. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 2 of 6)
7-35

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

PRESS ALT: 2000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

180

170

150

160

140

150

150
140

140

130
130

130
120

MAX
RANGE

120

110
100

80
MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

100
90

80
MAX END
AND R / C

100

40
12

22 23

24.5

GW~
1000 LB

22 23

24.5

GW~
1000 LB

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-10. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 3 of 6)
7-36

AA0542
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

PRESS ALT: 2000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170
160
150

150

140

170
160
150

130

140
130

140

120

130

MAX
RANGE

MAX
RANGE
120

110

MAX END
AND R / C

TES
100

80
MAX END
AND R / C

50
12 14

22 23

12 14

24.5

GW~
1000 LB

GW~
1000 LB
20

110

100

TRANSMISSION TORQUE LIMIT

100

120

TORQUE AVAILABLE ~ 30 MINU

110

TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

180

TRUE AIRSPEED ~ KTS

ATF = 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
~ CONTINUOUS

180

40
30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0543
SA

Figure 7-10. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 4 of 6)
7-37

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

PRESS ALT: 2000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160
150

ATF = 0.9

~ CONTINUOUS

170

150

140

130
140

180

ATF = 1.0

160

~ CONTINUOUS

ATF = 1.0

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170
160
150
140

120
MAX
RANGE

MAX
RANGE

130

120

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

110

TORQUE AVAILABLE ~ 30 MINUTES

100

TRANSMISSION TORQUE LIMIT

110

100

TRUE AIRSPEED ~ KTS

110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

130

40
50

50
30

GW~ 12 14
1000 LB

23
24.5

GW~
1000 LB

24.5

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-10. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 5 of 6)
7-38

AA0544
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T700 (2)

PRESS ALT: 2000 FT
50oC

60oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

150
140

150

140

130

MAX
RANGE

120
TRANSMISSION TORQUE LIMIT

100
110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

110

100

90
MAX END
AND R / C

170
160

120
MAX
RANGE

180

150
140
130
120

MAX END
AND R / C

110

100

TRUE AIRSPEED ~ KTS

160

TORQUE AVAILABLE ~ 30 MINUTES

~ CONTINUOUS

170

ATF = 0.9

180

TRANSMISSION TORQUE LIMIT

160

~ CONTINUOUS

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

50
60
40

50
12

24.5

GW~
1000 LB

24.5

GW~
1000 LB

30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0545
SA

Figure 7-10. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 6 of 6)
7-39

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 4000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170

160
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
150

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140

140
140

130

130
130

120

MAX
RANGE

120

120
110
110

100
100

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

50
40

16 18

GW ~
1000 LB

14 16

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-11. Cruise - Pressure Altitude 4,000 Feet (Sheet 1 of 6)
7-40

AA0455
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 4000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

130

100

MAX END
AND R / C

120

110

100

MAX END
AND R / C

60
60

TRUE AIRSPEED ~ KTS

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

MAX
RANGE

TRANSMISSION TORQUE LIMIT

120

130

MAX
RANGE

60
50

40
12

14 16

GW ~
1000 LB

20
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0454
SA

Figure 7-11. Cruise - Pressure Altitude 4,000 Feet (Sheet 2 of 6)
7-41

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 4000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

140
MAX
RANGE

130

120

110
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

100

90
MAX END
AND R / C

100

80
MAX END
AND R / C
70

120

110

100
TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ CONTINUOUS

MAX
RANGE

130

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

130

GW ~
1000 LB

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-11. Cruise - Pressure Altitude 4,000 Feet (Sheet 3 of 6)
7-42

AA0453
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 4000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

180
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

170

150

160

140

150

140

100

110

100

80
MAX END
AND R / C

MAX END
AND R / C

50
12
40

GW ~
1000 LB

40
12

120

110

100

TRANSMISSION TORQUE LIMIT

ATF= 1.0

TORQUE AVAILABLE ~ 30 MIN ATF= 0.9

110

130

120

~ CONTINUOUS

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

~ CONTINUOUS

130

140

130

~ 30 MIN ATF= 0.9

MAX
RANGE

GW ~
1000 LB
20

TRUE AIRSPEED ~ KTS

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

40
30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0456
SA

Figure 7-11. Cruise - Pressure Altitude 4,000 Feet (Sheet 4 of 6)
7-43

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 4000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

160

150

140

150

180
170
160
150

130

140

MAX
RANGE

130

120

110

140

MAX END
AND R / C
80

80
MAX END
AND R / C

110

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MINUTES

100

TRANSMISSION TORQUE LIMIT

110

120

100

TRUE AIRSPEED ~ KTS

130

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

ATF= 0.9

160

~ CONTINUOUS

ATF= 1.0

ATF= 0.9

170

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 1.0

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

60
40

50
12

40
12

22
40

GW ~
1000 LB

30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-11. Cruise - Pressure Altitude 4,000 Feet (Sheet 5 of 6)
7-44

AA0457
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 4000 FT
50OC

60OC

TOTAL FUEL FLOW ~ 100 LB/HR

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
180

170
160

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

~ CONTINUOUS

150

140

ATF = 1.0

ATF = 0.9

TOTAL FUEL FLOW ~ 100 LB/HR

~ CONTINUOUS

ATF = 1.0

ATF = 0.9

IAS ~ KTS

180
170
160

150
140

150

MAX
RANGE

130

MAX
RANGE

140

120

130

MAX END
AND R / C
80

MAX END
AND R / C

110

100

TRUE AIRSPEED ~ KTS

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE
~ 30 MINUTES

110

120
100
TRANSMISSION TORQUE LIMIT

120

110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

130

50
60

60
40

50
30

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0458
SA

Figure 7-11. Cruise - Pressure Altitude 4,000 Feet (Sheet 6 of 6)
7-45

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

PRESS ALT: 4000 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140
140

130

120

120
MAX
RANGE

110

80
MAX END
AND R / C
70

100
TRANSMISSION TORQUE LIMIT

100

TRANSMISSION TORQUE LIMIT

100

110

80
MAX END
AND R / C

14 16

24.5

GW ~
1000 LB
30

40
40

TRUE AIRSPEED ~ KTS

MAX
RANGE

110

TRUE AIRSPEED ~ KTS

130

14 16

24.5

GW ~
1000 LB
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-12. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 1 of 6)
7-46

AA0546
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

PRESS ALT: 4000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

180

170

150

160

140

150

150
140
140

130
130

130
120

120

MAX
RANGE

120

110
100

100

80
MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

80
MAX END
AND R / C

40
12

14 16

24.5

50
12

GW ~
1000 LB

24.5

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0547
SA

Figure 7-12. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 2 of 6)
7-47

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

PRESS ALT: 4000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

180

170
150
160
140

150

160
150

130

140
130

140

120

130

MAX
RANGE

120

MAX
RANGE
110

120

100

MAX END
AND R / C

80
MAX END
AND R / C

110

TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

170

14 16

24.5

GW ~
1000 LB

24.5

GW ~
1000 LB

100

TRUE AIRSPEED ~ KTS

180

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-12. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 3 of 6)
7-48

AA0548
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

PRESS ALT: 4000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

130

ATF = 1.0

140

ATF = 0.9

160

~ CONTINUOUS

170

ATF = 0.9

~ CONTINUOUS

180
150

170
160
150

140

140
120

130
120

120
100

110

90
MAX END
AND R / C

24.5

GW ~
1000 LB

40
30

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MINUTES

110

100

TRUE AIRSPEED ~ KTS

130

MAX
RANGE

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

MAX
RANGE

24.5

40
12
30

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0549
SA

Figure 7-12. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 4 of 6)
7-49

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

PRESS ALT: 4000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

~ CONTINUOUS

170
160
150

150

140

130

140

120
MAX
RANGE

130

180
170
160
150
140

MAX
RANGE

110

130

120

100
110

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

110
TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

ATF = 0.9

180

70
MAX END
AND R / C
60

24.5

24.5
50

GW ~
1000 LB

30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-12. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 5 of 6)
7-50

60
40

50
40

100

TRUE AIRSPEED ~ KTS

160

~ CONTINUOUS

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0550
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T700 (2)

PRESS ALT: 4000 FT
50oC

60oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

TRUE AIRSPEED ~ KTS

140
MAX
RANGE

130

150

130

120

ATF = 1.0

140

170
160
150
140

110
130

120

100

110

100

90
70
80
60
70

180

MAX END
AND R / C

120
110
MAX
RANGE

100

24.5

60
GW ~
1000 LB

GW ~
1000 LB

23
22

24.5

TRUE AIRSPEED ~ KTS

150

TORQUE AVAILABLE ~ 30 MINUTES

160

ATF = 0.9

~ CONTINUOUS

170

ATF = 0.9

180

~ CONTINUOUS

ATF = 1.0
TORQUE AVAILABLE ~ 30 MINUTES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

40
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0551
SA

Figure 7-12. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 6 of 6)
7-51

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 6000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

130

140

130

100

80
MAX END
AND R / C
70

110

100

TRUE AIRSPEED ~ KTS

MAX END
AND R / C

OUS & 30 MIN

120

110

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

110

MAX
RANGE

120

TRANSMISSION TORQUE LIMIT

MAX
RANGE

TORQUE AVAILABLE ~ CONTINU

120

40
40

GW ~
1000 LB

20
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-13. Cruise - Pressure Altitude 6,000 Feet (Sheet 1 of 6)
7-52

AA0461
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 6000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

140
130

130
MAX
RANGE

120

130

MAX
RANGE

120

MAX END
AND R / C

OS & 30 MIN

TORQUE AVAILABLE ~ CONTINU

TRANSMISSION TORQUE LIMIT

100

80
MAX END
AND R / C
70

40
12

GW~
1000 LB

100

70
TRANSMISSION TORQUE LIMIT

30 MIN

110
100

TRUE AIRSPEED ~ KTS

110

TORQUE AVAILABLE ~ CONTINUOUS &

TRUE AIRSPEED ~ KTS

110

40
30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0460
SA

Figure 7-13. Cruise - Pressure Altitude 6,000 Feet (Sheet 2 of 6)
7-53

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 6000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170
10

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150

170

150

160

140

150

130

140

140
MAX
RANGE

130

MAX
RANGE

120

130

120

50
12

GW ~
1000 LB

40
30

MAX END
AND R / C

ATF= 0.9
TORQUE AVAILABLE ~ 30 MIN

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

30 MIN ATF= 0.9

100

40
12

120

110

100

TRANSMISSION TORQUE LIMIT AFT~1.0

100

~ CONTINUOUS

110

TORQUE AVAILABLE ~

TRUE AIRSPEED ~ KTS

110

TRUE AIRSPEED ~ KTS

180

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-13. Cruise - Pressure Altitude 6,000 Feet (Sheet 3 of 6)
7-54

AA0459
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 6000 FT
10oC

20oC

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

ATF= 1.0

ATF= 0.9

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

160

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 1.0

ATF= 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

150
160
140

180
170
160

150

150
130

140
MAX
RANGE

120

130

120

MAX END
AND R / C
80

MAX END
AND R / C

110
TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ 30 MINUTES

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MIN

110

50
60
40
50
12

GW ~
1000 LB

22
40

GW ~
1000 LB

100

TRUE AIRSPEED ~ KTS

110

120

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

130

140

MAX
RANGE

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0462
SA

Figure 7-13. Cruise - Pressure Altitude 6,000 Feet (Sheet 4 of 6)
7-55

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 6000 FT

30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
20

150

140

150

130
MAX
RANGE

140

180
170
160
150

MAX
RANGE

120

140

130

110

TRUE AIRSPEED ~ KTS

120

120
100

MAX END
AND R / C

80
MAX END
AND R / C

110
TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ 30 MINUTES

100

TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ 30 MINUTES

110

100

TRUE AIRSPEED ~ KTS

160

ATF = 0.9

170

160

~ CONTINUOUS

ATF = 1.0

~ CONTINUOUS

ATF = 0.9

180

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

50
30

22
20

GW ~
1000 LB

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

TORQUE PER ENGINE ~ %

100
AA0463A
SA

Figure 7-13. Cruise - Pressure Altitude 6,000 Feet (Sheet 5 of 6)

7-56

Change 8

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 6000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

~ CONTINUOUS

180
170
160
150

MAX
RANGE

160

150

140

130

120

130

110

120

100

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

TORQUE AVAILABLE ~ 30 MINUTES

ATF = 0.9

ATF = 1.0

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

100

MAX END
AND R / C

60
40
50

22
30

GW~
1000 LB

30
20

10
0

0
20

100

TORQUE PER ENGINE ~ %

DATA BASIS: FLIGHT TEST

AA0464A
SA

Figure 7-13. Cruise - Pressure Altitude 6,000 Feet (Sheet 6 of 6)
7-57

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

PRESS ALT: 6000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

180

170

150

160

140

150

150
140
140
130
130
130
120
120

MAX END
AND R / C

OUS & 30 MIN
MAX END
AND R / C

100

40
12 14 16

12
30

24.5

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ CONTINU

100

TRANSMISSION TORQUE LIMIT

OUS & 30 MIN

110

MAX
RANGE

110

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

120
MAX
RANGE

14 16

24.5

GW ~
1000 LB

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-14. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 1 of 6)
7-58

AA0552
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

PRESS ALT: 6000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

180

170
170

150
160

160
140

150

130

120
MAX
RANGE

130
MAX
RANGE

MAX END
AND R / C

OUS & 30 MIN

100

120

100

MAX END
AND R / C

50
12

24.5

12
30

GW ~
1000 LB

TORQUE AVAILABLE ~ CONTINU

110

TRANSMISSION TORQUE LIMIT

OUS & 30 MIN

110

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

120

140

24.5

GW ~
1000 LB

110

100

TRUE AIRSPEED ~ KTS

130

TRANSMISSION TORQUE LIMIT

140

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0553
SA

Figure 7-14. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 2 of 6)
7-59

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

PRESS ALT: 6000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160
150

30
180

~ CONTINUO
US

170

150

140

130

170

ATF = 0.9

~ CONTINUOUS

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160
150

140

140
120

100

90
MAX END
AND R / C

50
23

24.5

40
12

100
110
90

MAX END
AND R / C

100

TRANSMISSION TORQUE LIMIT

110

120

TORQUE AVAILABLE ~ 30 MINUTES

MAX
RANGE

120

MAX
RANGE

24.5

GW ~
1000 LB

30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-14. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 3 of 6)
7-60

50
12

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

130
110

ATF = 0.9
TORQUE AVAILABLE ~
30 MINUTES
ATF = 1.0 TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

AA0554
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

PRESS ALT: 6000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160
150

180

150

140

130

140

170
160
150

120

130

MAX
RANGE

140
130

MAX
RANGE

110

120

100
110

110

60
MAX
END
AND
R/C

24.5

40
12

50
40

24.5

12
30

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

MAX
END
AND
R/C

TORQUE AVAILABLE ~ 30 MIN

TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MIN

TRUE AIRSPEED ~ KTS

GW ~
1000 LB

100

TRUE AIRSPEED ~ KTS

170

ATF = 1.0

~ CONTINUOUS

ATF = 0.9

180

ATF = 0.9

~ CONTINUOUS

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0555
SA

Figure 7-14. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 4 of 6)
7-61

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

PRESS ALT: 6000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

MAX
RANGE

130
120

120

110

MAX
RANGE

80
MAX
END
AND
R/C

24.5

TRANSMISSION TORQUE LIMIT

100

ATF = 1.0

ATF = 0.9

160

100

110

170

150
140
130
120
110

100

MAX END
AND R / C

24.5
60

40
GW ~
1000 LB 12

30
GW ~
1000 LB

40
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASIS: FLIGHT TEST

Figure 7-14. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 5 of 6)
7-62

TRUE AIRSPEED ~ KTS

140

130

180

TORQUE AVAILABLE ~ 30 MIN

150

TORQUE AVAILABLE ~ 30 MIN

160

140

~ CONTINUOUS

ATF = 0.9

~ CONTINUOUS

170

ATF = 1.0

150

180

TRUE AIRSPEED ~ KTS

AA0556A
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T700 (2)

PRESS ALT: 6000 FT
50OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
11

12
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

ATF = 0.9

180

160
150

TRUE AIRSPEED ~ KTS

140
MAX
RANGE

130

140

130

TORQUE AVAILABLE ~ 30 MINUTES

~ CONTINUOUS

170

ATF = 1.0

150

120
110

120

110

100

100
90

60
MAX END
AND R / C

40
GW ~
1000 LB

40
20
30
20

10
0

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0557
SA

Figure 7-14. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 6 of 6)
7-63

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

130

140

130

50
12

14 16

80
MAX END
AND R / C
70

50
12

GW ~
1000 LB

100
TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

110

TORQUE AVAILABLE ~ CON

100

TINUOUS & 30 MIN

OUS & 30 MIN

110

120

110

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

MAX
RANGE

TRUE AIRSPEED ~ KTS

120

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-15. Cruise - Pressure Altitude 8,000 Feet (Sheet 1 of 6)
7-64

AA0467
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170

180
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

170

150

160

140

150

130

140

150

140

130

80
MAX END
AND R / C

120

110

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

90
TRANSMISSION TORQUE LIMIT

100

TORQUE AVAILABLE ~

100

CONTINUOUS & 30 MIN

TINUOUS & 30 MIN

110

TORQUE AVAILABLE ~ CON

TRUE AIRSPEED ~ KTS

120

130

MAX
RANGE

12
40

GW ~
1000 LB

100

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0466
SA

Figure 7-15. Cruise - Pressure Altitude 8,000 Feet (Sheet 2 of 6)
7-65

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
11

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

IAS ~ KTS
12

170

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

160

ATF= 1.0

ATF= 0.9

~ CONTINUOUS

ATF= 0.9

ATF= 1.0

~ CONTINUOUS

180
170
160

140
150
150
130

140

140
MAX
RANGE

MAX END
AND R / C
80

50
12
40

TORQUE AVAILABLE ~ 30 MIN

100

80
MAX END
AND R / C

GW ~
1000 LB

110

100

TRANSMISSION TORQUE LIMIT

100

TRANSMISSION TORQUE LIMIT

110

120

TRUE AIRSPEED ~ KTS

130

110

120

TORQUE AVAILABLE ~ 30 MIN

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

130

50
40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-15. Cruise - Pressure Altitude 8,000 Feet (Sheet 3 of 6)
7-66

AA0465
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

150

160

180
170

140

150

160

130

150

140
120

MAX
RANGE

140

MAX
RANGE

130

90
MAX END
AND R / C
80

MAX END
AND R / C

70
50

110
TRANSMISSION TORQUE LIMIT

100

120

100

TORQUE AVAILABLE ~ 30 MINUTES

110

TRANSMISSION TORQUE LIMIT

120

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

110

100

TRUE AIRSPEED ~ KTS

170

160

ATF= 1.0

ATF= 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 1.0

180

ATF= 0.9

~ CONTINUOUS

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

50
12
40

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0468
SA

Figure 7-15. Cruise - Pressure Altitude 8,000 Feet (Sheet 4 of 6)
7-67

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
5

150

TRUE AIRSPEED ~ KTS

140

MAX
RANGE

140

MINUTES

150

ATF= 1.0

ATF= 0.9

180
170

130

120
MAX
RANGE

160
150
140

130

110

120

100

110

100

120

MAX END
AND R / C

130

MAX END
AND R / C

50
60

40
50

40
30

GW ~
1000 LB

30
20

0
10

TORQUE PER ENGINE ~ %

DATA BASIS: FLIGHT TEST

Figure 7-15. Cruise - Pressure Altitude 8,000 Feet (Sheet 5 of 6)
7-68

TRUE AIRSPEED ~ KTS

160

TORQUE AVAILABLE ~ 30

170

160

~ CONTINUOUS

180

TES

ATF= 1.0

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TORQUE AVAILABLE ~ 30 MINU

ATF= 0.9

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0469A
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
11

ATF = 1.0

ATF = 0.9

~ CONTINUOUS

180
170
160
150

TORQUE AVAILABLE ~ 30 MINUTES

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

150

140

130

120
MAX
RANGE

140

TRUE AIRSPEED ~ KTS

110
130
100
120
90
110
80

100

90
MAX END
AND R / C

80
70

40
12

22
30

GW~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0470
SA

Figure 7-15. Cruise - Pressure Altitude 8,000 Feet (Sheet 6 of 6)
7-69

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

PRESS ALT: 8000 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

180

170
170

150
160

160
140

150

120

110

MAX END
AND R / C

50
12
40

TORQUE AVAILABLE

100

MAX
RANGE

100

TRANSMISSION TORQUE LIMIT

OUS & 30 MIN

MAX
RANGE

110

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

120

MAX END
AND R / C

130

120

110

50
12

GW ~
1000 LB

24.5
40

GW ~
1000 LB

24.5

100

TRUE AIRSPEED ~ KTS

130

140

~ CONTINUOUS & 30

130

TRANSMISSION TORQUE LIMIT

140

MIN

150

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-16. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 1 of 6)
7-70

AA0558
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

PRESS ALT: 8000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

30
180

150

170

170
160

140

160

150

130
140

140
120
130

90
MAX
END
AND
R/C

MIN

100

~ CONTINUOUS & 30

100

120

80
MAX
END
AND
R/C

110
TRANSMISSION TORQUE LIMIT

110

MAX
RANGE

TORQUE AVAILABLE

OUS & 30 MIN

MAX
RANGE

TRANSMISSION TORQUE LIMIT

120

100

TRUE AIRSPEED ~ KTS

110

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

130

40
50

12 14

24.5
30

GW ~
1000 LB

24.5

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0559
SA

Figure 7-16. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 2 of 6)
7-71

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

PRESS ALT: 8000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

150

140

130

140

110

100

110
MAX
RANGE

100

170

TORQUE AVAILABLE

TRANSMISSION TORQUE LIMIT

MAX
RANGE

120

180

160

120

130

150
140
130
120
110

100
80
90
70

MAX END
AND R / C

80
60

24.5

GW ~
1000 LB

40
30

MAX END
AND R / C
TRANSMISSION TORQUE LIMIT

TES

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

~ 30 MINUTES

160

ATF = 1.0

170

ATF = 0.9

~ CONTINUOUS

180

50
24.5
40
12

30
GW ~
1000 LB
20

50
40
30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-16. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 3 of 6)
7-72

TRUE AIRSPEED ~ KTS

160

ATF = 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0560
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

PRESS ALT: 8000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

120

ATF = 1.0

110
MAX
RANGE
100

110

170
160
150
140
130
120

100

80
MAX
END
AND
R/C

TRANSMISSION TORQUE LIMIT

60
MAX
END
AND
R/C

60
12

50
40

24.5

12
30

GW ~
1000 LB

100
90

40
23

110

TRUE AIRSPEED ~ KTS

MAX
RANGE

180

TORQUE AVAILABLE ~ 30 MIN

120

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140
130

130

TORQUE AVAILABLE

150

ATF = 0.9

160

140

~ CONTINUOUS

ATF = 0.9

~ CONTINUOUS

170

~ 30 MIN ATF = 1.0

150
180

24.5
50

GW ~
1000 LB

20
30

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0561
SA

Figure 7-16. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 4 of 6)
7-73

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

PRESS ALT: 8000 FT
30OC

40OC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

TRUE AIRSPEED ~ KTS

140
130

MAX
RANGE

120
110

120

ATF = 1.0

160

110
MAX
RANGE

100

100
90

60
MAX END
AND R / C

150
140
130
120
110
100
90
80

MAX END
AND R / C

170

TRUE AIRSPEED ~ KTS

150

130

180

TORQUE AVAILABLE ~ 30 MINUTES

160

ATF = 0.9

~ CONTINUOUS

170

140

~ CONTINUOUS

ATF = 0.9

180

TORQUE AVAILABLE ~ 30 MINUTES ATF = 1.0

150

60
23

GW ~
1000 LB

40
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-16. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 5 of 6)
7-74

AA0562
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T700 (2)

PRESS ALT: 8000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
11

12
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

150
140

MAX
RANGE

130

120

TORQUE AVAILABLE ~ 30 MINUTES

160

TRUE AIRSPEED ~ KTS

ATF = 0.9

~ CONTINUOUS

170

ATF = 1.0

140

180

120
110
100

110

100

70
90
60
80
MAX END
AND R / C

50
23
40

60
12

50
GW ~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0563
SA

Figure 7-16. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 6 of 6)
7-75

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 10000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

180

170

120

MAX
RANGE

130

120

100

90
MAX END
AND R / C

MAX
RANGE

110

TRANSMISSION TORQUE LIMIT

110

170
160
150

140

130

100

120

110

80
MAX END
AND R / C

50
12

GW ~
1000 LB

100

TRUE AIRSPEED ~ KTS

130

140

TORQUE AVAILAB

140

TRANSMISSION TORQUE LIMIT

150

LE ~ CONTINUOU

160

OUS & 30 MIN.

ABLE ~ CONTINU

TORQUE AVAIL

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

S & 30 MIN.

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

TRUE AIRSPEED ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-17. Cruise - Pressure Altitude 10,000 Feet (Sheet 1 of 5)
7-76

AA0420
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 10000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

160

140

MAX
RANGE

130

120

110

100

90
MAX END
AND R / C

MAX
RANGE

110

100
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

TORQUE AVAIL

130

170
160
150
140

80
MAX END
AND R / C

130

120

110
TRANSMISSION TORQUE LIMIT

150

180

150

ABLE ~ CONTINU

160

MIN

170

~ CONTINUOUS & 30

TORQUE AVAILABLE

OUS & 30 MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

50
60

100

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

60
40

50
12

GW ~
1000 LB

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0419
SA

Figure 7-17. Cruise - Pressure Altitude 10,000 Feet (Sheet 2 of 5)
7-77

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 10000 FT
−10oC

0 oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

MAX
RANGE

140

130

120
MAX
RANGE

110

120

ATF=1.0

180
170
160
150
140
130

120

100

90
MAX END
AND R / C

110

TRANSMISSION TORQUE LIMIT

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

150

70
MAX END
AND R / C
60

50
60

100

TRUE AIRSPEED ~ KTS

140

TORQUE AVAILABLE ~ 30 MINUTES

150

ATF=0.9

ATF=1.0

160

~ CONTINUOUS

170

160

TES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

180

ATF=0.9

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

50
12
40

GW ~
1000 LB

20
10
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-17. Cruise - Pressure Altitude 10,000 Feet (Sheet 3 of 5)
7-78

AA0418
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 10000 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

MAX
RANGE

130

120

120
MAX
RANGE
110

ATF = 1.0

170
160
150
140
130

100

110

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

180

90
80
MAX END
AND R / C
70

120

110

100

60
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

140

150

TORQUE AVAILABLE

150

TRANSMISSION TORQUE LIMIT

160

TES

170

~ 30 MINUTES

180

160

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

ATF = 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 0.9

170

80
70

60
50

30
40

12
20

GW ~
1000 LB

30
20

GW ~
1000 LB

30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0421
SA

Figure 7-17. Cruise - Pressure Altitude 10,000 Feet (Sheet 4 of 5)
7-79

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 10000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

MAX
RANGE

130

120
MAX
RANGE

110
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

120

110

100
90

100

MAX END
AND R / C

ATF= 1.0

170
160
150
140
130
120

70
MAX END
AND R / C

180

60
12

GW ~
1000 LB

30
20

80
70

40
30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-17. Cruise - Pressure Altitude 10,000 Feet (Sheet 5 of 5)
7-80

100

110

TRUE AIRSPEED ~ KTS

150

TES

160

TRANSMISSION TORQUE LIMIT

170

150

TORQUE AVAILABLE ~ 30 MINU

180

160

~ CONTINUOUS

TORQUE AVAILABLE ~ 30 MINU

ATF= 1.0

ATF= 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 0.9

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0422
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

PRESS ALT: 10000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

150

130

120
MAX
RANGE

120

110
MAX
RANGE

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

110

140

TORQUE AVAIL

140

100

MAX
END
AND
R/C

180
170
160
150
140
130
120

110

70
MAX
END
AND
R/C

TRANSMISSION TORQUE LIMIT

ABLE ~ CONTINU

160

150

OUS

170

100

TRUE AIRSPEED ~ KTS

160

ABLE ~ CONTINU

TORQUE AVAIL

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

OUS & 30 MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

12 14

24.5

12
30

GW ~
1000 LB

24.5

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0564
SA

Figure 7-18. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 1 of 5)
7-81

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

PRESS ALT: 10000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

160

OUS & 30 MIN

170

150
140

140

MAX
RANGE

120

130

120

TORQUE AVAIL

110

110
MAX
RANGE

100

180
170
160
150
140
130
120
110

100

MAX
END
AND
R/C

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

150

ABLE ~ CONTINU

160

OUS & 30 MIN

180

ABLE ~ CONTINU

TORQUE AVAIL

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
MAX
END
AND
R/C

40
24.5

50
12
40

24.5

GW ~
1000 LB

50
40
30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-18. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 2 of 5)
7-82

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0565
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

PRESS ALT: 10000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140
110

100

MAX
RANGE

60
MAX
END
AND
R/C

30
GW ~
1000 LB

100
90

GW ~
1000 LB

110

120

70
24.5

24.5

130

MAX
END
AND
R/C

ATF = 1.0

ATF = 0.9

~ CONTINUOUS

150

TRUE AIRSPEED ~ KTS

100

160

TRANSMISSION TORQUE LIMIT

110

170

MINUTES

MAX
RANGE

120

180

TORQUE AVAILABLE ~ 30

130

TORQUE AVAI

TRUE AIRSPEED ~ KTS

140

120

ATF = 1.0
LABLE ~ 30 MI

150

130

TRANSMISSION TORQUE LIMIT

160

140

NUTES

170

ATF = 0.9

~ CONTINUOUS

180

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0566
SA

Figure 7-18. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 3 of 5)
7-83

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

PRESS ALT: 10000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

TRUE AIRSPEED ~ KTS

140
130

MAX
RANGE

120
110
100

110

ATF = 1.0

ATF = 0.9

160

MAX
RANGE

100

170

150

140
130
120
110
100

TRUE AIRSPEED ~ KTS

120

180

TORQUE AVAILABLE ~ 30 MIN

150

130

~ CONTINUOUS

160

140

TORQUE AVAILABLE ~ 30 MIN

170

ATF = 1.0

~ CONTINUOUS

ATF = 0.9

180

90
60

80
MAX END
AND R / C

70
23

60
GW ~
1000 LB 12

40
14

GW ~
1000 LB 12

60
14

22
50

22
40

40
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-18. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 4 of 5)
7-84

AA0567
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T700 (2)

PRESS ALT: 10000 FT
30OC

40OC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

MAX
RANGE

120
110

110

100

ATF = 1.0

ATF = 0.9

170
160

MAX
RANGE

100

150
140
130
120
110
100

70
90

TRUE AIRSPEED ~ KTS

130

120

180

TORQUE AVAILABLE ~ 30 MINUTES

140

~ CONTINUOUS

150

130

TORQUE AVAILABLE ~ 30 MINUTES

160

TRUE AIRSPEED ~ KTS

ATF = 0.9

~ CONTINUOUS

170

ATF = 1.0

140

180

90
60

80
MAX END
AND R / C

MAX END
AND R / C

60
GW ~
1000 LB

22
GW ~
1000 LB

20
50
40

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0568
SA

Figure 7-18. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 5 of 5)
7-85

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 12000 FT

−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

140

130

MIN

140

130

120

110

180
170
160
150

MAX
RANGE

140

130

120

100

110

110
100

90
MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

MAX
RANGE

150

TORQUE AVAILABLE

150

70
MAX END
AND R / C
60

100

TRUE AIRSPEED ~ KTS

160

120

160

~ CONTINUOUS & 30

TORQUE AVAILABLE

~ CONTINUOUS & 30

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

60
40

50
12

GW ~
1000 LB

GW ~
1000 LB
40

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %
AA0425
SA

Figure 7-19. Cruise - Pressure Altitude 12,000 Feet (Sheet 1 of 5)
7-86

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 12000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

150

130

120

MAX
RANGE
120

180
170
160
150
140

130

110
MAX
RANGE

120

100

110

100

90
MAX END
AND R / C

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

110

90
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

TORQUE AVAIL

140

150

ABLE ~ CONTINU

160

100

TRUE AIRSPEED ~ KTS

170

ABLE ~ 30 MIN

160

TORQUE AVAIL

OUS & 30 MIN

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

50
12

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0424
SA

Figure 7-19. Cruise - Pressure Altitude 12,000 Feet (Sheet 2 of 5)
7-87

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 12000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

100

90
80
MAX END
AND R / C

170
160
150
140
130

120

110

100

90
80

MAX END
AND R / C

40
50

GW ~
1000 LB

30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-19. Cruise - Pressure Altitude 12,000 Feet (Sheet 3 of 5)
7-88

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110

ATF = 1.0
~ 30 MINUTES

MAX
RANGE

100
TRANSMISSION TORQUE LIMIT

120

ATF = 0.9

120

110

180

TRANSMISSION TORQUE LIMIT

MAX
RANGE

130

TORQUE AVAIL

TRUE AIRSPEED ~ KTS

140

150

130

150

ABLE ~ 30 MINU

160

TORQUE AVAILABLE

170

160

~ CONTINUOUS

180

ATF = 1.0

TES

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0423
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 12000 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

130
120

MAX
RANGE

110

100

110

100
90
80

MAX END
AND R / C

TES

120

170
160
150
140
130

60
MAX END
AND R / C

120

110
100
90
80
70

60
12

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

MAX
RANGE

180

TRANSMISSION TORQUE LIMIT

130

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

TORQUE AVAILABLE

150

TORQUE AVAILABLE ~ 30 MINU

160

150

~ 30 MINUTES

170

ATF = 0.9

~ CONTINUOUS

180

ATF = 1.0

160

~ CONTINUOUS

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

GW ~
1000 LB

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0426
SA

Figure 7-19. Cruise - Pressure Altitude 12,000 Feet (Sheet 4 of 5)
7-89

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 12000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

180

ATF = 0.9

~ CONTINUOUS

170
160

ATF = 1.0
TORQUE AVAILABLE ~ 30 MINU
TES

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

150

140

130

120

TRUE AIRSPEED ~ KTS

150
140

110

MAX
RANGE

130

100

120

110
80
100
70
90
60
80
MAX END
AND R / C

60
50
12
40

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %

DATA BASIS: FLIGHT TEST

Figure 7-19. Cruise - Pressure Altitude 12,000 Feet (Sheet 5 of 5)
7-90

AA0427A
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

PRESS ALT: 12000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
4

170

160

OUS & 30 MIN

180
170

150
140

120
MAX
RANGE

110

140

130

120

TORQUE AVAIL

TRUE AIRSPEED ~ KTS

130

150

ABLE ~ CONTINU

160

OUS & 30 MIN

180
170
160

ABLE ~ CONTINU

110

100

150
140

TORQUE AVAIL

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MAX
RANGE

130
120
110

90
100

100
80

90
70

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

80
60

70
MAX END
AND R / C

MAX END
AND R / C

60
GW ~
1000 LB

40
12

GW ~
1000 LB

12 14

30
40

40
20

30
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0569
SA

Figure 7-20. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 1 of 5)
7-91

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

PRESS ALT: 12000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

180
140

130
MAX
RANGE

120

110

TORQUE AVAIL

140

130

110

100

MAX END
AND R / C

130
120

100

80
MAX END
AND R / C

23
14

140

110

150

MAX
RANGE

60
GW ~
1000 LB

160

100

170

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

150

TRUE AIRSPEED ~ KTS

150

ABLE ~ CONTINU

160

ABLE ~ 30 MIN

170

~ CONTINUOUS

180

TORQUE AVAIL

OUS & 30 MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

22
50

30
40

40
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-20. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 2 of 5)
7-92

AA0570
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

PRESS ALT: 12000 FT
−10OC

0 OC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
20

160

150

120

110
100
90

140

100
MAX
RANGE
90

80
MAX END
AND R / C

ATF = 1.0

ATF = 0.9

150

120
110
100
90
80

MAX END
AND R / C

130

MINUTES

MAX
RANGE

160

40
GW ~
1000 LB

TRUE AIRSPEED ~ KTS

LE ~ 30 MINUTES

130

170

110

TORQUE AVAILAB

TRUE AIRSPEED ~ KTS

140

120

130

180

TORQUE AVAILABLE ~ 30

150

~ CONTINUOUS

160

140

ATF = 1.0

170

ATF = 0.9

~ CONTINUOUS

180

60
GW ~
1000 LB

30
20

0
10

TORQUE PER ENGINE ~ %

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0571
SA

Figure 7-20. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 3 of 5)
7-93

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

PRESS ALT: 12000 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

150

110

TORQUE AVAILABLE ~ 30 MINUTES

140
130
MAX
RANGE
120
110
100
90
80

MAX
RANGE
90

160
150
140
130
120
110
100

70
90
60
80

MAX END
AND R / C

ATF = 0.9

100

170

MAX END
AND R / C

70
22

22
GW ~
1000 LB

GW ~
1000 LB

50
40

0
10

TORQUE PER ENGINE ~ %

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-20. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 4 of 5)
7-94

TRUE AIRSPEED ~ KTS

120

180

TORQUE AVAILABLE ~ 30 MINUTES ATF = 1.0

150

130

~ CONTINUOUS

160

TRUE AIRSPEED ~ KTS

ATF = 0.9

~ CONTINUOUS

170

ATF = 1.0

140

180

AA0572
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T700 (2)

PRESS ALT: 12000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160
30

150

~ CONTINUOUS

170

150
140

120

110

MAX
RANGE

130

TORQUE AVAILABLE ~ 30 MINUTES

160

TRUE AIRSPEED ~ KTS

ATF = 0.9

180

ATF = 1.0

140

120
110
100
90
80

100

MAX END
AND R / C

40
60
12
50

GW ~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0573
SA

Figure 7-20. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 5 of 5)
7-95

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 14000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160
150

120

MAX
RANGE

130

120

110
MAX
RANGE

100

180
170
160
150
140

130

120

110

MAX END
AND R / C

110

60
MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

140

TORQUE AVAIL

140

150

ABLE ~ CONTINU

170

160

100

TRUE AIRSPEED ~ KTS

OUS & 30 MIN

ABLE ~ CONTINU

TORQUE AVAIL

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

OUS & 30 MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

22
50

30
40

40
GW ~
1000 LB

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-21. Cruise - Pressure Altitude 14,000 Feet (Sheet 1 of 5)
7-96

AA0430
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 14000 FT

−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

150
140
130
MAX
RANGE

150

140

TORQUE AVAILABLE

160

130

120

110

180
170
160
150
140
130

MAX
RANGE
100

110

100

MAX END
AND R / C

120
90

TRANSMISSION TORQUE LIMIT

120

MAX END
AND R / C

50
14

GW ~
1000 LB

90
80
70

50
40

100

110

TRUE AIRSPEED ~ KTS

180

TRUE AIRSPEED ~ KTS

160

TORQUE AVAI
LABLE ~ 30 MI
NUTES

~ CONTINUOUS

~ CONTINUOUS & 30

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

30
20

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %
AA0429
SA

Figure 7-21. Cruise - Pressure Altitude 14,000 Feet (Sheet 2 of 5)
7-97

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 14000 FT

−10oC

0 oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

140

ATF = 0.9

ATF = 1.0

130

120

180
170
160
150
140

130

MAX
RANGE

130

100

110

100
90
80

MAX END
AND R / C

50
12
40

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

120
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110
MAX
RANGE

MAX END
AND R / C

120

110
100

TRUE AIRSPEED ~ KTS

150

140

TORQUE AVAI

160

~ CONTINUOUS

170

150

LABLE ~ 30 MI

ATF = 0.9

180

NUTES

160

ATF = 1.0

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TORQUE AVAILABLE

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ 30 MINUTES

170

90
80
70
60
50

30
12

20
40

GW ~
1000 LB

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %
AA0428
SA

Figure 7-21. Cruise - Pressure Altitude 14,000 Feet (Sheet 3 of 5)
7-98

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 14000 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
4

170

140
130
120

120

110
MAX
RANGE
100

110
100
90
80

12
60

160
150
140
130

120

110
100

90
60
MAX END
AND R / C

80
70

20
60

GW ~
1000 LB

GW ~
1000
LB

170

MAX END
AND R / C

TES

MAX
RANGE

130

180

TRUE AIRSPEED ~ KTS

150

140

TRANSMISSION TORQUE LIMIT

160

150

TORQUE AVAILABLE ~ 30 MINU

170

TORQUE AVAILABLE ~ 30 MINU

ATF = 0.9
~ CONTINUOUS

180

ATF = 0.9

160

~ CONTINUOUS

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

40
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0431
SA

Figure 7-21. Cruise - Pressure Altitude 14,000 Feet (Sheet 4 of 5)
7-99

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 14000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

ATF = 0.9

~ CONTINUOUS

180
170
160

ATF = 1.0
TORQUE AVAILABLE ~ 30 MINU
TES

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

150

140

130

120

150

TRUE AIRSPEED ~ KTS

110
140

MAX
RANGE

100

130
90
120
110

100

60
GW ~
1000 LB

50
70
40
60
12

MAX END
AND R / C

30
20

10
0

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-21. Cruise - Pressure Altitude 14,000 Feet (Sheet 5 of 5)
7-100

AA0432
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

PRESS ALT: 14000 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

180

180
170

140

170

160
160

130
150

150

120
140

100

OUS & 30 MIN

120
MAX
RANGE

110

100

MAX END
AND R / C

GW ~
1000 LB

MAX
RANGE

60
MAX END
AND R / C

130
120
110
100
90

40
22

TRUE AIRSPEED ~ KTS

TORQUE AVAILABLE ~ CON
TINUOUS & 30 MIN

130

TORQUE AVAILABLE ~ CONTINU

TRUE AIRSPEED ~ KTS

140
110

GW ~
1000 LB

22
12

40
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0574
SA

Figure 7-22. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 1 of 5)
7-101

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

PRESS ALT: 14000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

130
MAX
RANGE

120

110

TORQUE AVAIL

TRUE AIRSPEED ~ KTS

140

120

110

100
MAX
RANGE

ABLE ~ 30 MIN

150

180

130

ABLE ~ CONTINU

160

140

170
160
150

TORQUE AVAIL

170

~ CONTINUOUS

180

140
130
120
110

80
100

100
70

TRUE AIRSPEED ~ KTS

OUS & 30 MIN

150

90
60

80
MAX END
AND R / C

GW ~
1000 LB

80
MAX END
AND R / C

40
12

GW ~
1000 LB

22
12

20
50

40
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-22. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 2 of 5)
7-102

AA0575
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

PRESS ALT: 14000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
20

160

150

110
100

ATF = 1.0

160

MINUTES

ATF = 0.9

100
MAX
RANGE

170

150
140

130
120
110
100

70
90

TRUE AIRSPEED ~ KTS

MAX
RANGE

120

110

180

TORQUE AVAILABLE ~ 30

130

~ CONTINUOUS

140

120

TORQUE AVAILAB

TRUE AIRSPEED ~ KTS

150

130

LE ~ 30 MINUTES

160

ATF = 0.9

~ CONTINUOUS

170

ATF = 1.0

140
180

90
60

80
50

MAX END
AND R / C

60
GW ~
1000 LB

GW ~
1000 LB

20
50
40

0
10

TORQUE PER ENGINE ~ %

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0576
SA

Figure 7-22. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 3 of 5)
7-103

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

PRESS ALT: 14000 FT
10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

100
130

MAX
RANGE

180
170
160
150
140

MAX
RANGE

130

120

120
80

100
90
80
MAX END
AND R / C
70
GW ~
1000 LB

110

TORQUE AVAILABLE ~ 30 MIN

110

MAX END
AND R / C

GW ~
1000 LB

50
40

100
90
80
70
20
60

50
40

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-22. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 4 of 5)
7-104

TRUE AIRSPEED ~ KTS

140

110

TORQUE AVAILABLE ~ 30 MIN

TRUE AIRSPEED ~ KTS

150

120

ATF = 0.9

160

130

ATF = 1.0

~ CONTINUOUS

170

~ CONTINUOUS

ATF = 0.9

180

ATF = 1.0

140

AA0577
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T700 (2)

PRESS ALT: 14000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

10
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

140

150

110

140
MAX
RANGE

130

120

TORQUE AVAILABLE ~ 30 MINUTES

160

TRUE AIRSPEED ~ KTS

ATF = 0.9

~ CONTINUOUS

170

ATF = 1.0

130

180

120
110
100
90

100

MAX END
AND R / C

40
60

GW ~
1000 LB
20

40
30

20
10

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0578
SA

Figure 7-22. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 5 of 5)
7-105

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 16000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR

MIN

170

150

TRUE AIRSPEED ~ KTS

140
130

130

120

110
MAX
RANGE

100

110

100

140

MAX
RANGE

120

150

TORQUE AVAILABLE

160

180
170
160
150
140
130

120

110

100

90
80

MAX END
AND R / C

40
50

GW ~
1000 LB

40
30
20

0
10

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-23. Cruise - Pressure Altitude 16,000 Feet (Sheet 1 of 5)
7-106

TRUE AIRSPEED ~ KTS

180

160

MIN

~ CONTINUOUS & 30

TORQUE AVAILABLE

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ CONTINUOUS & 30

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0435
SA

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 16000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR

170

TRUE AIRSPEED ~ KTS

140
130

MAX
RANGE

120

TES

130

TORQUE AVAILABLE

150

140

120

110

180
170
160
150
140

MAX
RANGE

100

130

90
110
80
100
70
90
60

80
MAX END
AND R / C

120

110
100
90
80

MAX END
AND R / C

40
12

12
30

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

160

TRANSMISSION TORQUE LIMIT

170

150

~ CONTINUOUS & 30

180

ABLE ~ 30 MINU

160

~ CONTINUOUS

MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TORQUE AVAIL

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

GW ~
1000 LB

0
10

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASIS: FLIGHT TEST

AA0434A
SA

Figure 7-23. Cruise - Pressure Altitude 16,000 Feet (Sheet 2 of 5)
7-107

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 16000 FT

−10oC

0 oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

130

110

MAX
RANGE

100

TRANSMISSION TORQUE LIMIT

120
110
100
90
80
70

MINUTES

ATF = 1.0

120

MAX
RANGE

130

180
170
160
150
140
130

MAX END
AND R / C

120
110
100

TRUE AIRSPEED ~ KTS

140

~ CONTINUOUS

140

TRANSMISSION TORQUE LIMIT

150

TORQUE AVAILABLE ~ 30

160

TRUE AIRSPEED ~ KTS

LE ~ 30 MINUTES

170

150

TORQUE AVAILAB

180

ATF = 0.9

~ CONTINUOUS

ATF = 0.9

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
80
70

60
30

50
12

20
20

GW ~
1000 LB

50
12

20
40

GW ~
1000 LB

30
20

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %
AA0433
SA

Figure 7-23. Cruise - Pressure Altitude 16,000 Feet (Sheet 3 of 5)
7-108

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 16000 FT

10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR

170

160

150

140

130

120

150

110

140

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

MAX
RANGE

180
170
160
150

MAX
RANGE

100

140

TORQUE LIMIT

130
120
110
100

TRANSMISSION

TRUE AIRSPEED ~ KTS

170

90
80

130
90
120
80
110
70

100
90

60
MAX END
AND R / C

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

180

ATF = 0.9

~ CONTINUOUS

ATF = 1.0
TORQUE AVAILABLE
~ 30 MINUTES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

ATF = 0.9

TOTAL FUEL FLOW ~ 100 LB/HR

~ CONTINUOUS

TES

ATF = 1.0
TORQUE AVAILABLE ~ 30 MINU

IAS ~ KTS

70
40

60
12

18
50

GW ~
1000 LB

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %
AA0436
SA

Figure 7-23. Cruise - Pressure Altitude 16,000 Feet (Sheet 4 of 5)
7-109

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 16000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

12
170

180
170

TRUE AIRSPEED ~ KTS

160

ATF = 1.0

ATF = 0.9

160

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

TES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

140

130

120

110

150

100

140
MAX
RANGE

130

120
80
110
70
100
60

90
MAX END
AND R / C

70
40
60

18
30

GW~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %
DATA BASE: FLIGHT TEST

Figure 7-23. Cruise - Pressure Altitude 16,000 Feet (Sheet 5 of 5)
7-110

AA0437
SA

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

PRESS ALT: 16000 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150
180
140

180

170
170
130

160

MAX
RANGE

100
90

OUS & 30 MIN

110

100

MAX
RANGE

80
MAX END
AND R / C

130
120
110
100
90
80

MAX END
AND R / C

140

40
GW ~
1000 LB

TRUE AIRSPEED ~ KTS

120

150

TORQUE AVAILABLE ~ CONTINU

130

110

120

TORQUE AVAILABLE ~ CONTINU

140

TRUE AIRSPEED ~ KTS

160

OUS & 30 MIN

150

60
GW ~
1000 LB

40
20

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0579
SA

Figure 7-24. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 1 of 4)
7-111

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

PRESS ALT: 16000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140
180
180

160

130

MAX
RANGE

110
100
90

110

100

160
150
140
130

TORQUE AVAILABLE ~ 30

140

120

TORQUE AVAILABLE ~ CON
TINUOUS & 30 MIN

TRUE AIRSPEED ~ KTS

150

170

MINUTES

~ CONTINUOUS

130

MAX
RANGE

120
110
100
90

80
50

MAX END
AND R / C

20
GW ~
1000 LB

20
12

GW ~
1000 LB

18
50
40

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-24. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 2 of 4)
7-112

TRUE AIRSPEED ~ KTS

170

AA0580
SA

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

PRESS ALT: 16000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

110
100

MAX
RANGE

ATF = 1.0

ATF = 0.9

160

~ 30 MINUTES

100

170

150
140
130
120
110
100
90

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

MAX
RANGE

120

110

180

TORQUE AVAILABLE

140
130

120

~ CONTINUOUS

150

130

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

160

ATF = 1.0

~ CONTINUOUS

170

~ 30 MINUTES

180

ATF = 0.9

140

MAX END
AND R / C

70
40

GW ~
1000 LB

50
40

0
0

TORQUE PER ENGINE ~ %

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0581
SA

Figure 7-24. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 3 of 4)
7-113

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T700 (2)

PRESS ALT: 16000 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

140

MAX
RANGE

120
110
100

100

ATF = 0.9

MAX
RANGE

170
160
150
140
130
120
110

70
100
60

90
80

MAX END
AND R / C

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-24. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 4 of 4)
7-114

TRUE AIRSPEED ~ KTS

130

110

180

ATF = 1.0
TORQUE AVAILABLE ~ 30 MINUTES

140

~ CONTINUOUS

150

120

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

160

ATF = 0.9

~ CONTINUOUS

170

130
ATF = 1.0

180

AA0582
SA

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 18000 FT

−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS
9

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRUE AIRSPEED ~ KTS

130
MAX
RANGE

110

~ CONTINUOUS & 30

140

120

130

120

TORQUE AVAILABLE

150

180

110

100

170
160
150

TORQUE AVAILABLE

160

140

MAX
RANGE

140
130
120

110
80

100

100
70

TRUE AIRSPEED ~ KTS

~ CONTINUOUS & 30

170

MIN

150
180

MAX END
AND R / C

30
12

GW ~
1000 LB

18
40

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

TORQUE PER ENGINE ~ %
AA0440
SA

Figure 7-25. Cruise - Pressure Altitude 18,000 Feet (Sheet 1 of 5)
7-115

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 18000 FT

−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS
9

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

140
130

MAX
RANGE

TES

130

120

TORQUE AVAIL

150

TRUE AIRSPEED ~ KTS

140

110

100

120

180
170
160
150
140

MAX
RANGE

130

110

120

110

100

100
90

TRUE AIRSPEED ~ KTS

160

ABLE ~ 30 MINU

170

150

ABLE ~ CONTINU

180

TORQUE AVAIL

~ CONTINUOUS

OUS & 30 MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
60

20
80

80
MAX END
AND R / C

MAX END
AND R / C

70
40

60
12

GW ~
1000 LB

20
30

0
0

TORQUE PER ENGINE ~ %

TORQUE PER ENGINE ~ %
AA0439
SA

Figure 7-25. Cruise - Pressure Altitude 18,000 Feet (Sheet 2 of 5)
7-116

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 18000 FT

−10oC

0 oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
10

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

MAX
RANGE

120
110

ATF = 0.9

60
MAX END
AND R / C

MAX
RANGE

100

ATF = 1.0

100

170
160
150
140
130
120
110

TRUE AIRSPEED ~ KTS

140
130

110

TORQUE AVAILABLE ~ 30

TRUE AIRSPEED ~ KTS

150

120

180

TORQUE AVAILABLE ~ 30 MINUTES

160

130

~ CONTINUOUS

170

140

MINUTES

180

ATF = 0.9

~ CONTINUOUS

150

100
90
MAX END
AND R / C

70
40

60
12

16
50

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

AA0438
SA

Figure 7-25. Cruise - Pressure Altitude 18,000 Feet (Sheet 3 of 5)
7-117

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 18000 FT

10oC

20oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
9

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

130
120

ATF = 0.9

100

170
160
150
140
130
120

110

110
70

100

TRUE AIRSPEED ~ KTS

TES

140

110

180

TORQUE AVAILABLE ~ 30 MINUTES

150

~ CONTINUOUS

160

120

TORQUE AVAILABLE ~ 30 MINU

170

130

ATF = 1.0

ATF = 0.9
ATF = 1.0
~ CONTINUOUS

180

TRUE AIRSPEED ~ KTS

140

100
60

MAX
RANGE

MAX END
AND R / C

MAX
RANGE

MAX END
AND R / C

40
60

60
30

GW ~
1000 LB

10
0

0
0

TORQUE PER ENGINE ~ %

0
0

TORQUE PER ENGINE ~ %
AA0441
SA

Figure 7-25. Cruise - Pressure Altitude 18,000 Feet (Sheet 4 of 5)
7-118

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 18000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

10
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

ATF = 0.9

ATF = 1.0

TES

180
170

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

150

140

130

120

TRUE AIRSPEED ~ KTS

160

110

150
100
140

MAX
RANGE

130
80

120
110

100
60
90
MAX END
AND R / C

70
12
60

GW~
1000 LB

30
20

10
0

0
0

TORQUE PER ENGINE ~ %

DATA BASIS: FLIGHT TEST

AA0442A
SA

Figure 7-25. Cruise - Pressure Altitude 18,000 Feet (Sheet 5 of 5)
7-119

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

PRESS ALT: 18000 FT
−50OC

−40OC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

110

130
120
MAX
RANGE

180
170
160
150

100

90
MAX
RANGE

140
130
120
110

100

70
90

TRUE AIRSPEED ~ KTS

140

110

120

TORQUE AVAILABLE

150

~ CONTINUOUS & 30

160

130

TORQUE AVAILABLE

170

TRUE AIRSPEED ~ KTS

140

~ CONTINUOUS & 30

180

MIN

150

90
60

80
50

MAX END
AND R / C

60
30

50
12

GW ~
1000 LB

30
20

10
0

0
0

TORQUE PER ENGINE ~ %

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-26. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 1 of 4)
7-120

AA0583
SA

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

PRESS ALT: 18000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

MIN

140

TORQUE AVAILABLE

140
130
MAX
RANGE

120

110

~ 30 MIN

170
160

100

90
MAX
RANGE

150
140
130
120

80
110

100

100
90

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

150

120

TORQUE AVAILABLE

160

180
~ CONTINUOUS

170

TRUE AIRSPEED ~ KTS

130

~ CONTINUOUS & 30

180

MAX END
AND R / C

70
GW ~
1000 LB

40
12

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0584
SA

Figure 7-26. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 2 of 4)
7-121

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

PRESS ALT: 18000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140

MAX
RANGE

110

90
MAX
RANGE
80

ATF = 0.9

ATF = 1.0

160

~ 30 MIN

100

170

150
140
130
120
110

100

TRUE AIRSPEED ~ KTS

130
120

110

180

TORQUE AVAILABLE

140

~ CONTINUOUS

150

~ 30 MIN

TRUE AIRSPEED ~ KTS

160

120

TORQUE AVAILABLE

~ CONTINUOUS

170

ATF = 0.9
ATF = 1.0

130
180

100
60

90
80

MAX END
AND R / C

70
GW ~
1000 LB

40
12

MAX END
AND R / C
GW ~
1000 LB

30
50

30
10

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-26. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 3 of 4)
7-122

AA0585
SA

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T700 (2)

PRESS ALT: 18000 FT
10OC

20OC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
10

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

140

MAX
RANGE

120
110
100

ATF = 1.0

ATF = 0.9

160

90
MAX
RANGE

170

150

60
90

140
130
120
110

TRUE AIRSPEED ~ KTS

130

100

180

TORQUE AVAILABLE ~ 30 MINUTES

140

~ CONTINUOUS

150

110

TES

160

ATF = 0.9

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

170

120

TORQUE AVAILABLE ~ 30 MINU

180

ATF = 1.0

130

100
90

MAX END
AND R / C

40
GW ~
1000 LB

70
GW ~
1000 LB

50
20

30
10

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0586
SA

Figure 7-26. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 4 of 4)
7-123

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 20000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
4

160

140

130

120

110

MIN

180
170
160
150
140

100

130

130
MAX
RANGE

110

TRANSMISSION TORQUE LIMIT

MAX
RANGE

120

80
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

150

100
90
80
MAX END
AND R / C

70
60
50

GW ~
1000 LB

60
MAX END
AND R / C
50

120
110
100

TRUE AIRSPEED ~ KTS

170

150

TORQUE AVAILABLE

180

160

~ CONTINUOUS & 30

TORQUE AVAILABLE

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ CONTINUOUS & 30

MIN

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
80
70

40
12

GW ~
1000 LB

40
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-27. Cruise - Pressure Altitude 20,000 Feet (Sheet 1 of 4)
7-124

AA0445
SA

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 20000 FT
−30oC

−20oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
2

170

140
130

130

120

110

180
170
160
150
140

100

120
110
100
90
MAX END
AND R / C

MAX
RANGE

TRANSMISSION TORQUE LIMIT

MAX
RANGE

TES

140

130
120

80
110
70

TRUE AIRSPEED ~ KTS

150

TORQUE AVAIL

160

150

TORQUE AVAILABLE

170

~ CONTINUOUS

180

TRUE AIRSPEED ~ KTS

160

MIN

~ CONTINUOUS & 30

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ABLE ~ 30 MINU

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

100

90
MAX END
AND R / C

70
40

GW ~
1000 LB

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0444
SA

Figure 7-27. Cruise - Pressure Altitude 20,000 Feet (Sheet 2 of 4)
7-125

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 20000 FT
−10oC

0oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
2

170

~ CONTINUOUS

170

TRUE AIRSPEED ~ KTS

160
150
140

150

140

130

120

110

100

180
170
160
150
140

MAX
RANGE

130

120

MAX
RANGE

130
120

110

70
100

100
60

90
MAX END
AND R / C

MAX END
AND R / C

80
70

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %
DATA BASE:

Change 8

TORQUE PER ENGINE ~ %

FLIGHT TEST

Figure 7-27. Cruise - Pressure Altitude 20,000 Feet (Sheet 3 of 4)

7-126

TRUE AIRSPEED ~ KTS

180

ATF= 0.9

~ CONTINUOUS

ATF= 0.9
ATF= 1.0
TORQUE AVAILABLE ~ 30 MINU
TES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

ATF= 1.0
TORQUE AVAILABLE ~ 30 MINUTES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

AA0443A
SA

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

CLEAN CONFIGURATION
PRESS ALT: 20000 FT
10oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
10
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

~ CONTINUOUS

ATF = 0.9
ATF = 1.0

140

180
170

120

110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

160
150
140
MAX
RANGE

130

120
110
100

100

60
90
MAX END
AND R / C

GW ~
1000 LB

40
30
20

10
0

0
0

TORQUE PER ENGINE ~ %

AA0446
SA

Figure 7-27. Cruise - Pressure Altitude 20,000 Feet (Sheet 4 of 4)
7-127

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

PRESS ALT: 20000 FT
−50oC

−40oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
2

170

160
150
140

MIN

140

130

180

120

110

100

170
160
150
140

130

130
90

120
MAX
RANGE

110

120

MAX
RANGE

110

100

100
90

80
70

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

170

150

TORQUE AVAILABLE

180

TRUE AIRSPEED ~ KTS

160

~ CONTINUOUS & 30

TORQUE AVAILABLE

~ CONTINUOUS & 30

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
MIN

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MAX END
AND R / C
70

40
60

60
GW ~
1000 LB

30
12

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-28. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 1 of 4)
7-128

AA0587
SA

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

PRESS ALT: 20000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

MIN

140

MAX
RANGE

110

90
MAX
RANGE

160
150
140

130
120
110

100

TRUE AIRSPEED ~ KTS

130

100

LE ~ 30 MINUTES

140

120

110

TORQUE AVAILABLE

150

170

TORQUE AVAILAB

160

180
120

~ CONTINUOUS

170

TRUE AIRSPEED ~ KTS

130

~ CONTINUOUS & 30

180

100
60

MAX END
AND R / C

GW ~
1000 LB

30
10

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA0588
SA

Figure 7-28. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 2 of 4)
7-129

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

PRESS ALT: 20000 FT
−10OC

0OC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

140

130
MAX
RANGE

120
110
100

ATF = 0.9

MAX
RANGE

170
160
150
140
130
120
110
100

60
90

90
50

MAX END
AND R / C

70
12

GW ~
1000 LB

30
10

0
0

TORQUE PER ENGINE ~ %

0
0

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7-28. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 3 of 4)
7-130

TRUE AIRSPEED ~ KTS

140

100

180

TORQUE AVAILABLE ~ 30
MINUTES

150

110

ATF = 1.0

160

120

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

170

TORQUE AVAILABLE ~ 30
MINUTES

ATF = 0.9

180

ATF = 1.0

130

AA0589
SA

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T700 (2)

PRESS ALT: 20000 FT
10OC

TOTAL FUEL FLOW ~ 100 LB/HR
3

10
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

140

130

160
150
140
130

MAX
RANGE

120
110
100

110

100

INDICATED AIRSPEED ~ KTS

ATF = 0.9

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

170

ATF = 1.0
TORQUE AVAILABLE ~ 30 MINUTES

120

180

90
50
MAX END
AND R / C

40
70
12

GW ~
1000 LB

40
30

20
10
0

0
0

TORQUE PER ENGINE ~ %
DATA BASE: FLIGHT TEST

AA0590
SA

Figure 7-28. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 4 of 4)
7-131

TM 1-1520-237-10

Section V OPTIMUM CRUISE
7.19 OPTIMUM RANGE CHARTS.
This section presents a method to optimize cruise performance for long range missions when the altitudes flown
are not restricted by other requirements. The optimum altitude for maximum range chart (Figure 7-29) provides the
pressure altitude at which to cruise to obtain the maximum
possible range for any gross weight and FAT conditions.

7-132

The altitude determined for optimum range may also be
used for optimum endurance. Enter the chart at a current
cruise or takeoff temperature condition and move along the
temperature guide lines to the anticipated gross weight for
cruise and obtain the optimum pressure altitude. Turn to the
cruise chart closest to the altitude and temperature predicted
by the optimum range chart for specific cruise information.
The use of this chart is shown by the example.

TM 1-1520-237-10

OPTIMUM RANGE
CLEAN CONFIGURATION
100% RPM R
HIRSS (BAFFLES INSTALLED)
EXAMPLE
WANTED:

METHOD:

CRUISE ALTITUDE FOR OPTIMUM RANGE
AND CORRESPONDING CRUISE CHART FOR
FLIGHT CONDITIONS

ENTER CHART AT FAT (24 OC), MOVE RIGHT
TO REFERENCE / OPTIMUM PRESSURE ALTITUDE
(1,500 FT). MOVE PARALLEL WITH THE
TEMPERATURE TREND LINES TO AIRCRAFT
GROSS WEIGHT (16,500 LB). MOVE LEFT OR
RIGHT PARALLELING THE TEMPERATURE TREND
LINE TO NEAREST EVEN THOUSAND
REFERENCE / OPTIMUM PRESSURE ALTITUDE
LINE (12,000). MOVE LEFT TO FREE AIR
TEMPERATURE LINE (2.5 OC), MOVE UP OR DOWN
TO NEAREST TEN VALUE ON THE FREE AIR
TEMPERATURE SCALE (0 OC).

KNOWN:
REFERENCE CONDITIONS OF:
PRESSURE ALTITUDE = 1,500 FT
FAT = 24 OC
GROSS WEIGHT = 16,500 LB

SELECT CRUISE CHART WITH ALTITUDE AND
TEMPERATURE DATA AT THE NEAREST
REFERENCE / OPTIMUM PRESSURE ALTITUDE
(12,000 FT) AND THE NEAREST TEN DEGREE
FREE AIR TEMPERATURE (0 OC).

22
21

GROSS WEIGHT
~ 1000 LBS

19
18

17
16
15

14
13

FREE AIR TEMPERATURE ~ o C

−10

−20

−30

TEM
PER
ATU
TRE
RE
ND
LIN
ES

−40

−50

−60
0

DATA BASIS:

FLIGHT TEST

OPTIMUM PRESSURE ALTITUDE ~ 1000 FT

AA0683_1B
SA

Figure 7-29. Optimum Altitude For Maximum Range (Sheet 1 of 2)

7-133

TM 1-1520-237-10

OPTIMUM RANGE
HIGH DRAG CONFIGURATION 100% RPM R
HIRSS (BAFFLES INSTALLED)

23
22

GROSS WEIGHT
~ 1000 LB

21
20

19
18
17
40

16
15
14

FREE AIR TEMPERATURE ~ OC

20
TEM
PER
ATU
TRE
RE
ND
LIN
ES

−10

−20

−30

−40

−50

−60
0

OPTIMUM PRESSURE ALTITUDE ~ 1000 FT

DATA BASIS:

FLIGHT TEST

Figure 7-29. Optimum Altitude For Maximum Range (Sheet 2 of 2)

7-134

AA0683_2B

TM 1-1520-237-10

Section VI DRAG
7.20 EXTERNAL LOAD DRAG CHART.
The general shapes of typical external loads are shown
on Figure 7-30 as a function of the load frontal area. The
frontal area is combined with the typical drag coefficient of
the general shapes to obtain a drag multiplying factor for
use with the 10 sq. ft. drag scale on each cruise chart. The
TRQ ˜% value obtained from the cruise chart is multiplied by the drag multiplying factor and added to indicated
torque to obtain total torque required at any airspeed.

7.21 AIRCRAFT
CHANGES FOR
CHARTS.

CONFIGURATION
USE WITH CLEAN

DRAG
CRUISE

When external equipment or configuration differs from
the baseline clean configuration as defined in Section I, a
drag correction should be made similarly to the external
drag load method. Typical configuration changes that have
drag areas established from flight test or analysis along with
their drag multiplying factor are shown on Table 7-1.

Table 7-1. Configuration Drag Change
DRAG CHANGES FOR USE WITH CLEAN CRUISE CHARTS
Change in Flat Plate Drag
Area- F Sq. Ft.

Drag Multiplying Factor

a. Both cargo doors open

6.0

0.60

b. Cargo doors removed

4.0

0.40

c. Cargo mirror installed

0.3

0.03

d. IR Countermeasure Transmitter (ALQ-144) installed

0.8

0.08

e. Chaff Dispenser Installed

0.3

0.03

f. HIRSS not installed

-2.2

-0.22

0.3

0.03

3.8

0.38

i. Blade Erosion Kit

2.0

0.20

j. Skis installed

3.0

0.30

Item

g. Flare Dispenser

h. EH-60A Mission Antennas Only

7.22 AIRCRAFT
CONFIGURATION
DRAG
CHANGES FOR USE WITH HIGH DRAG CRUISE
CHARTS.

nal drag load method. Typical high drag configuration
changes that have been established from flight test or analysis along with the drag multiplying factors are shown.

When external equipment differs from the baseline high
drag configuration as defined in this Section, a drag correction should be made using Figure 7-31 similar to the exter-

7-135

TM 1-1520-237-10

EXTERNAL DRAG LOAD METHOD:

EXAMPLE

KNOWN:

WANTED:

ENTER CHART AT SYMBOL
FOR CYLINDER
MOVE RIGHT TO 80 SQ FT.
MOVE DOWN AND READ
DRAG MULTIPLYING FACTOR
= 4.5

SHAPE OF EXTERNAL
LOAD = CYLINDER
FRONTAL AREA OF
EXTERNAL LOAD = 80 SQ FT

DRAG MULTIPLYING FACTOR
DUE TO EXTERNAL LOAD

LOAD
DRAG

INCREASE IN DRAG AREA DUE TO EXTERNAL LOAD
0

100

120

140

160

180

200

220

240

SPHERE

STREAMLINED
CYLINDER

CYLINDER
CUBE
FLAT
PLATE
BOX

100

FRONTAL AREA
OF EXTERNAL
LOAD ~ SQ FT

BOX
IN
NET
0

DATA BASIS:

ESTIMATED

Figure 7-30. External Load Drag

7-136

DRAG MULTIPLYING FACTOR

24
AA0684A
SA

TM 1-1520-237-10

DRAG CONFIGURATIONS

CHANGE
IN
FLAT
PLATE
DRAG
F
SQ FT

DRAG
MULTI−
PLYING
FACTOR

ESSS − CLEAN, PYLONS REMOVED

−4.0

−0.40

ESSS − FOUR PYLONS / NO STORES

−1.7

−0.17

ESSS−TWO 450−GALLON TANKS INBOARD

0.5

0.05

−TWO 230−GALLON TANKS INBOARD

0.0

0.00

2.5

0.25

2.0

0.20

32.5

3.25

10.5

1.05

SKIS INSTALLED

3.0

0.30

BOTH CARGO DOORS OPEN

6.0

0.60

BOTH CARGO DOORS REMOVED

4.0

0.40

CARGO MIRROR INSTALLED

0.3

0.03

IR COUNTERMEASURE TRANSMITTER (ALQ−144) REMOVED

−0.8

−0.08

CHAFF DISPENSER REMOVED

−0.3

−0.03

HIGH DRAG
CRUISE CHART BASELINE
SPECIAL MISSION EQUIPMENT CONFIGURATIONS

ESSS−TWO 230−GALLON TANKS OUTBOARD
−TWO 450−GALLON −TANKS INBOARD

ESSS − FOUR 230−GALLON TANKS

VOLCANO SYSTEM INSTALLED (BOTH RACKS)
* VOLCANO CORRECTION MUST BE USED WITH HIGH DRAG CHARTS ONLY
VOLCANO SYSTEM INSTALLED (LOWER RACKS ONLY)

AA0685B
SA

Figure 7-31. Typical High Drag Configurations
Change 8

7-137

TM 1-1520-237-10

Section VII CLIMB - DESCENT
7.23 CLIMB/DESCENT CHART.
The CLIMB/DESCENT chart (Figures 7-32 and 7-33)
presents the rate of climb or descent resulting from an increase or decrease of engine torque from the value required
for level flight above 40 KIAS. The data are presented at
100% RPM R for various gross weights. The charts may
also be used in reverse to obtain the torque increase or

7-138

reduction required to achieve a desired steady rate of climb
or descent. The maximum R/C may be determined by subtracting the cruise chart torque required from the maximum
torque available at the desired flight conditions. Then enter
the difference on the torque increase scale of the climb
chart, move up to the gross weight, and read the resulting
maximum R/C.

TM 1-1520-237-10

CLIMB/DESCENT
CLEAN CONFIGURATION
100% RPM R
FOR IAS ABOVE 40 KIAS
4000

GROSS
WEIGHT
~ 1000 LB

DESCENT

14
16

3500

EXAMPLE
WANTED:
INDICATED TORQUE CHANGE FOR
DESIRED RATE−OF−CLIMB OR DESCENT.

KNOWN:
GROSS WEIGHT = 18,000 POUNDS
DESIRED RATE = 550 FEET PER MINUTE

RATE OF DESCENT ~ FT/MIN

20
22

3000

2500

2000

1500

1000

METHOD:
ENTER CHART AT 550 FEET PER MINUTE
MOVE RIGHT TO INTERSECT GROSS
WEIGHT LINE. MOVE DOWN TO READ
12% TRQ CHANGE.

500

0
0

TORQUE REDUCTION PER ENGINE ~ % TRQ
3500

GROSS
WEIGHT
~ 1000 LB

CLIMB

14
16

3000

RATE OF CLIMB ~ FT/MIN

20
22
2500

2000

1500

1000

500

0
0

TORQUE INCREASE PER ENGINE ~ % TRQ
DATA BASIS:

FLIGHT TEST

80
AA0687A
SA

Figure 7-32. Climb/Descent

7-139

TM 1-1520-237-10

CLIMB/DESCENT
100% RPM R
AIRSPEEDS ABOVE 40 KIAS
4000

GROSS
WEIGHT
~ 1000 LB

DESCENT

16
18

3500

RATE OF DESCENT ~ FT/MIN

20
22

3000

24
2500

2000

1500

1000

500

0
0

TORQUE REDUCTION PER ENGINE ~ % TRQ
3500

GROSS
WEIGHT
~ 1000 LB

12
CLIMB

14
16

RATE OF CLIMB ~ FT/MIN

3000

18
20

2500

22
24

2000

1500

1000

500

0
0

TORQUE INCREASE PER ENGINE ~ % TRQ
DATA BASIS:

FLIGHT TEST

AA0688A
SA

Figure 7-33. Climb/Descent - High Drag

7-140

TM 1-1520-237-10

Section VIII FUEL FLOW
7.24 IDLE FUEL FLOW.

Dual-engine idle fuel flow is presented as a function of
altitude at 0°C FAT in Table 7-2. The data are based on
operation at 62% to 69% Ng for idle and 85% to 89% for
flat pitch (collective full down) at 100% RPM R. Fuel flow
for the auxiliary power unit (APU) is also presented for a
nominal load of 80% maximum power as a function of
altitude and 0°C FAT for general planning.
7.25 SINGLE-ENGINE FUEL FLOW.
a. Engine fuel flow is presented in Figure 7-34 for various torque and pressure altitudes at a baseline FAT of 0°C
with engine bleed air extraction off. When operating at
other than 0°C FAT, engine fuel flow is increased 1% for
each 20°C above the baseline temperature and, decreased
1% for each 20°C below the baseline temperature.
b. To determine single-engine fuel flow during cruise,
enter the fuel flow chart at double the torque required for
dual-engine cruise as determined from the cruise charts and
obtain fuel flow from the single-engine scale. The singleengine torque may not exceed the transmission limit shown
on the chart. With bleed air on, single-engine fuel flow
increases as follows:

(2) When the IR suppressor system is installed and
operating in the benign mode (exhaust baffles removed),
the single-engine fuel flow will decrease about 8 lbs/hr.
7.26 DUAL-ENGINE FUEL FLOW.
Dual-engine fuel flow for level flight is presented on the
cruise charts in Section IV. For other conditions dualengine fuel flow may be obtained from Figure 7-34 when
each engine is indicating approximately the same torque by
averaging the indicated torques and reading fuel flow from
the dual-engine fuel flow scale. When operating at other
than the 0° FAT baseline, dual-engine fuel flow is increased
1% for each 20°C above baseline and is decreased 1% for
each 20°C below baseline temperature. With bleed air on,
dual-engine fuel flow increases as follows:
a. With bleed-air extracted, fuel flow increases:
(1) Engine anti-ice on -About 60 lbs/hr
Example: (760 lbs/hour = 820 lbs/hr).
(2) Heater on - About 20 lbs/hr
(3) Both on - About 80 lbs/hr

(1) With bleed-air extracted, fuel flow increases:
(a) Engine anti-ice on - About 30 lbs/hr
(b) Heater on - About 10 lbs/hr

b. When the cruise or hover IR suppressor system is
installed and operating in the benign mode (exhaust baffles
removed), the dual-engine fuel flow will decrease about 16
lbs/hr.

Table 7–2. Dual Engine Idle and Auxiliary Power Unit Fuel Flow
Pressure Altitude Feet

Ng = 62-69%
Ground Idle (No Load)
Lb/Hr

Ng = 85-89%
Flat Pitch (100% RPM R)
Lb/Hr

APU (Nominal) Lb/Hr

350

580

120

4,000

326

500

105

8,000

268

440

12,000

234

380

16,000

206

320

20,000

182

270

7-141

TM 1-1520-237-10

SINGLE/DUAL−ENGINE FUEL FLOW
100% RPM R
FAT = 0oC
BLEED AIR OFF
HIRSS (BAFFLES INSTALLED)
DUAL−ENGINE FUEL FLOW ~ LB/HR
200

400

600

800

1000

1200

1400

1600

115
TRANSMISSION LIMIT − 1 ENGINE

110

SL
105

100

95
4
90

INDICATED TORQUE PER ENGINE ~ %

85
8

70
12
65

60
16

45
20

PRESSURE ALTITUDE
~ 1000 FT

NOTE
INCREASE FUEL FLOW
1% FOR EACH 20 oC
ABOVE 0 oC FAT AND
DECREASE FUEL FLOW
1% FOR EACH 20 oC
BELOW 0 oC FAT.

20 16 12 8

20
100

200

SL
300

400

500

SINGLE−ENGINE FUEL FLOW ~ LB/HR

700

800
AA0689
SA

Figure 7-34. Single/Dual-Engine Fuel Flow

7-142

600

TM 1-1520-237-10

Section IX AIRSPEED SYSTEM CHARACTERISTICS
7.27 AIRSPEED SYSTEM CHARACTERISTICS
NOTE
Indicated airspeeds below 40 KIAS are unreliable. Airspeed conversion data KIAS to
KTAS for speeds above 40 KIAS are provided in Section IV CRUISE.
There are two different pitot-static systems on the UH60A. The type of airspeed system may be determined from
the mounting of the pitot-static probes on the cabin roof.
The pitot-static probes were originally flush mounted, the
modified pitot-static probes are mounted on a wedge which
rotates the pitot tube 20° further outboard and 3° nose
down. The wedge is covered by an aerodynamic fairing to
prevent ice accretion. The modified airspeed system was
used to derive the IAS presented on the charts in this
manual.
7.28 AIRSPEED CHARTS.
7.28.1 Airspeed Correction Charts. All indicated airspeeds shown on the cruise charts are based on level flight
of an aircraft with wedge mounted pitot static probes. Figures 7-35 through 7-37 provide the airspeed correction to
be added to the cruise chart IAS values to determine the
related airspeed indicator reading for other than level flight
mode. There are small variations in airspeed system errors
in level flight between those aircraft with wedge mounted
pitot static probes and those without wedge mounted pitot
static probes. Correction for these small variations is not
normally warranted. There are relatively large variations in
airspeed system error associated with climbs and descents.

These errors are significant and Figures 7-35 through 7-37
are provided primarily to show the general magnitude and
direction of the errors associated with the various flight
modes. If desired, these figures may be used in the manner
shown by the examples to calculate specific airspeed corrections.
7.28.2 Airspeed System Dynamic Characteristics.
The dynamic characteristics of the pilot and copilot airspeed indicating systems are normally satisfactory. However, the following anomalies in the airspeed and IVSI indicating system may be observed during the following
maneuvers or conditions:
a. During takeoffs, in the speed range of 40 to 80 KIAS,
5 to 10 KIAS airspeed fluctuation may be observed on the
pilot’s and copilot’s airspeed indicators.
b. Power changes in high power, low airspeed climbs
may cause as much as 30 knot airspeed changes in indicated airspeed. Increase in power causes increase in indicated airspeed, and a decrease in power causes decrease in
indicated airspeed.
c. The pilot and copilot airspeed indicators may be unreliable during high power climbs at low airspeeds (less
than 50 KIAS) with the copilot system reading as much as
30 knots lower than the pilot system.
d. On aircraft with wedge mounted pitot static probes
kit, in-flight opening and closing of doors and windows
may cause momentary fluctuations of approximately 300
feet per minute on the vertical speed indicators.

7-143

TM 1-1520-237-10

AIRSPEED SYSTEM CORRECTION
CLEAN
AIRCRAFT WITHOUT
WEDGE MOUNTED
PITOT−STATIC PROBES

CORRECTION TO ADD ~ KNOTS

15
R / C GREATER THAN 1400 FT / MIN
10

AUTOROTATION

5
DIVE
0

−5
LEVEL FLIGHT
−10
R / C LESS THAN 1400 FT / MIN
−15
20

100

120

140

160

180

IAS FROM CRUISE ~ KNOTS

EXAMPLE
WANTED:
INDICATED AIRSPEED TO OBTAIN MAX
RANGE FOR AN AIRCRAFT WITHOUT WEDGE
MOUNTED PITOT−STATIC PROBES.

KNOWN:
125 KIAS FOR MAX RANGE CRUISE
CHART AT A GIVEN PRESSURE ALTITUDE,
FAT, AND GROSS WEIGHT.

METHOD:
ENTER AT KNOWN IAS FROM CRUISE
CHART FOR MAX RANGE, MOVE UP TO LEVEL FLIGHT
LINE, MOVE LEFT, READ CORRECTION TO ADD TO
IAS = −2 KIAS. CALCULATE IAS FOR MAX RANGE:
= 125 KIAS − 2 KIAS = 123 KIAS.

DATA BASIS:

FLIGHT TEST

Figure 7-35. Airspeed Correction Aircraft Without Wedge Mounted Pitot-Static Probes

7-144

AA0690C
SA

TM 1-1520-237-10

AIRSPEED SYSTEM CORRECTION
CLEAN
WITH WEDGE MOUNTED PITOT−STATIC PROBES

CORRECTION TO ADD ~ KNOTS

20
R / C GREATER THAN 1400 FT / MIN

10
AUTOROTATION

DIVE
LEVEL
FLIGHT

−5
R / C LESS THAN 1400 FT / MIN
−10

−15
20

100

120

140

160

180

IAS FROM CRUISE CHARTS ~ KNOTS

EXAMPLE
WANTED:
INDICATED AIRSPEED TO CLIMB AT
MAXIMUM RATE OF CLIMB FOR AN
AIRCRAFT WITH WEDGE MOUNTED
PITOT−STATIC PROBES.

KNOWN:
70 KIAS MAX END / AND R / C FROM
APPROPRIATE CRUISE CHART FOR
A GIVEN PRESSURE ALTITUDE, FAT,
AND GROSS WEIGHT.

METHOD:
ENTER AT KNOWN IAS FROM
CRUISE CHART, MOVE UP TO R / C GREATER
THAN 1400 FPM, MOVE LEFT READ CORRECTION
TO ADD TO IAS = + 12.5 KTS, RE−ENTER
AT KNOWN IAS FROM CRUISE CHART, MOVE UP
TO R / C LESS THAN 1400 FPM LINE, MOVE LEFT,
READ CORRECTION TO ADD TO IAS = − 4 KTS
CALCULATE IAS FOR MAX R / C WHEN:
FOR R / C GREATER THAN 1400 FPM, AIRSPEED = 70 KIAS + 12.5 KIAS = 82.5 KIAS
FOR R / C LESS THAN 1400 FPM, AIRSPEED = 70 KIAS − 4 KIAS = 66 KIAS

DATA BASIS:

FLIGHT TEST

AA0691B
SA

Figure 7-36. Airspeed Correction Aircraft With Wedge Mounted Pitot-Static Probes

7-145

TM 1-1520-237-10

AIRSPEED SYSTEM CORRECTION

CORRECTION TO BE ADDED ~ KNOTS

10
AUTOROTATION
5
LEVEL FLIGHT
0
R / C GREATER THAN 1400 FT / MIN
−5

−10
R / C LESS THAN 1400 FT / MIN
−15

−20
20

100

120

140

160

IAS FROM HIGH DRAG CRUISE CHARTS ~ KNOTS

DATA BASIS: FLIGHT TEST

AA1029A
SA

Figure 7-37. Airspeed Correction Chart - High Drag

7-146

TM 1-1520-237-10

Section X SPECIAL MISSION PERFORMANCE
7.29 SPECIAL MISSION FLIGHT PROFILES.
Figures 7-38 through 7-40 show special mission flight
profiles required to obtain near maximum range when
equipped with the ESSS in three different tank configurations. The upper segment of each chart provides the recommended altitude profile along with the IAS and average
TRQ versus distance traveled. An average value of elapsed
time is also presented on the lower axis of the altitude
scale. The lower segment of each chart provides the relationship between fuel remaining and distance traveled resulting from the flight profile shown. This portion may be
utilized to check actual inflight range data to provide assurance that adequate range is being achieved. The chart is
divided into 3 regions of Adequate Range, Inadequate
range-return to base, and Inadequate range-requiring emergency action. When an inflight range point is in the Adequate range region, the required mission range can be obtained by staying on the recommended flight profile.
However, the range may not be achieved if stronger headwinds are encountered as the flight progresses, and normal
pilot judgement must be used. These charts also assume
that the flight track is within proper navigational limits.
Standard temperature variation with PA is shown on the
upper segment of the charts. A general correction for temperature variation is to decrease IAS by 2.5 KTS and total
distance traveled by 0.5% for each 10°C above standard.
Detailed flight planning must always be made for the actual
aircraft configuration, fuel load, and flight conditions when
maximum range is required. This data is based on JP-4 fuel.
It can be used with JP-5 or JP-8, aviation gasoline, or any
other approved fuels ONLY IF THE TAKEOFF GROSS
WEIGHT AND THE FUEL LOAD WEIGHT MATCH
THE DATA AT THE TOP OF THE CHART. The Flight
Time and the Distance Traveled data SHOULD NOT be
used with any full tank configuration if the fuel density is
not approximately 6.5 lb/gal (JP-4 fuel).
a. SELF-DEPLOYMENT MISSION. The selfdeployment mission is shown in Figure 7-38 and the ESSS
is configured with two 230-gallon tanks outboard and two
450-gallon tanks inboard. In this configuration, the aircraft
holds in excess of 11,000 lb of JP4 fuel and has a take-off
gross weight of 24,500 pounds in order to achieve the desired mission range of 1,150 Nm. This gross weight is allowed for ferry missions only, requiring low load factors

and less than 30 degree angle banked turns. This mission
was calculated for a standard day with a constant 10 knot
headwind added to be conservative. Since there may not be
any emergency landing areas available, the mission should
not be attempted if headwinds in excess of 10 knots are
forecast. Take-off must be made with a minimum of fuel
used (60 pounds) for engine start and warm-up, and a climb
to 2,000 feet should be made with maximum power and
airspeed between 80 and 105 KIAS. The first segment
should be maintained at 2,000 feet and 105 KIAS for 2
hours. The average engine TRQ should be about 79% for
this segment, but will initially be a little more and gradually
decrease. Altitude is increased in 2,000 feet increments to
maintain the optimum altitude for maximum range to account for fuel burn. The first 2 segments are for 2 hours
each, followed by 1 hour segments until reaching 10,000
feet. At this altitude, the airspeed for best range should also
be reduced to 95 KIAS for the remainder of the flight.
Engine bleed air was assumed to be off for this mission
except for that required for fuel tank pressurization. Electrical cabin heat may be used. Removal of the HIRSS
baffles (benign mode) will reduce fuel flow by about 16
lbs/hr. If oxygen is available, continuation of the staircase
climb sequence to 15,500 feet PA will result in about 23
additional Nm of range capability.
b. ASSAULT MISSION PROFILE - 4 tanks. The assault mission profile is shown in Figure 7-39 with the ESSS
configured with four 230-gallon tanks. In this configuration, the aircraft holds in excess of 8,300 pounds of JP4
fuel and assumes a take-off gross weight of 22,000 pounds
which provides a maximum mission range of 1140 Nm
with 400 lbs reserve. This mission was calculated for a
standard day with a zero headwind. Take-off must be made
with a minimum of fuel used (80 pounds) for engine start
and warm-up, and a Climb to 4,000 feet should be made
with maximum power and airspeed between 80 and 108
KIAS. The first segment should be maintained at 4,000 feet
and 108 KIAS for 1 hour. The average engine TRQ should
be about 79% for this segment, but will initially be a little
more and gradually decrease. Altitude is increased in 2,000
feet increments to maintain the optimum altitude for maximum range to account for fuel burn. The segments are for 1
hour each, until reaching 10,000 feet. At this altitude, the
airspeed for best range should be reduced to 95 KIAS for
the remainder of the flight.

7-147

TM 1-1520-237-10

c. ASSAULT MISSION PROFILE - 2 tanks. The assault mission profile is shown in Figure 7-40 with the ESSS
configured with two 230-gallon tanks. In this configuration,
the aircraft holds in excess of 5,300 pounds of JP4 fuel and
assumes a take-off gross weight of 22,000 pounds which
provides a maximum mission range of 630 Nm. with 400 lb
reserve. This mission was calculated for a standard day
with a zero headwind. Take-off must be made with a minimum of fuel used (80 lbs) for engine start and warm-up,

and a Climb to 4,000 feet should be made with max power
and airspeed between 80 and 108 KIAS. The first segment
should be maintained at 4,000 feet and 108 KIAS for 1
hour. The average engine TRQ should be about 77% for
this segment, but will initially be a little more and gradually
decrease as shown on each segment. Altitude is increased
in 2,000 feet increments to maintain the optimum altitude
for maximum range to account for fuel burn. At this altitude, the airspeed for best range should also be reduced to
95 KIAS for the remainder of the flight.

EXAMPLE:
WANTED:
Assurance of adequate aircraft range for mission defined.
KNOWN:
Flight position: 300 nm from base
Flight Track Within Limits
Fuel Remaining = 7,800 pounds
Elapsed flight time = 2 HRS, 50 MINS (2.83 HRS)
Target: Normal Flight Conditions:
Airspeed = 105 KIAS
Press Alt = 4,000 feet
Approx Torque = 75%
METHOD:
(1)

Enter chart at total distance flown and at fuel remaining, move to intersection and plot point. If point
falls on or above fuel remaining line (adequate range), remaining fuel is adequate to complete the
mission. If point falls below the fuel remaining line in the inadequate range, abort mission region,
immediately return to departure point while continuing to utilize altitide profile using total elapsed
flight time (see item 2). If point falls below the fuel remaining line in the inadequate range, region,
consult emergency procedures for corrective action.

(2)

To determine target nominial flight conditions, enter upper chart at elapsed flight time and move up
to determine target airspeed, approximate torque, and pressure altitude.
Figure 7-38. Self Deployment Mission Profile (Sheet 1 of 2)

7-148

Change 2

TM 1-1520-237-10

SELF DEPLOYMENT MISSION PROFILE

105 KIAS

PRESSURE ALT ~ 1000 FT

−9

95 KIAS
−5

(RECOMMENDED AIRSPEEDS)
(~62%)

(~52%)

−1
(~73%)

3
(~72%)

(APPROX TRQ~%)
(~75%)

STANDARD FAT ~ oC

ESSS/2−230 AND 2−450 GALLON TANK CONFIGURATION
STANDARD DAY 10 KT HEADWIND
HIRSS SUPPRESSED MODE
GROSS WEIGHT = 24,500 LB FUEL LOAD = 11,000 LB (JP4)
BLEED AIR OFF
60 LB WARM UP

11
(~80%)

15
0

NOMINAL FLIGHT TIME ~ HRS
12000

11000

(APPROX FULL FUEL − JP4)

10000

9000

8000
(7800 LBS)
DESIRED MISSION RANGE

FUEL REMAINING ~ LBS

7000
ADEQUATE
RANGE
FU
EL

6000

INADEQUATE RANGE
ABORT MISSION

5000

RE
M
AI
NI
NG

4000

3000

INADEQUATE
RANGE

2000

1000
LOW FUEL LIGHTS

DATA BASIS:

0
FLIGHT TEST

200

400

600

800

1000

1200

DISTANCE TRAVELED ~ NM

AA0608_2B
SA

Figure 7-38. Self Deployment Mission Profile (Sheet 2 of 2)
Change 2

7-149

TM 1-1520-237-10

ASSAULT MISSION PROFILE FOR MAX RANGE

(RECOMMENDED AIRSPEEDS)
108
KIAS

102
KIAS

100
KIAS

95 KIAS

PRESSURE ALT ~ 1000 FT

−9

−5
(~65%)

(~51%)

−1
(~70%)

3
(APPROX TRQ ~ %)

(~71%)
4

7
(~78%)

STANDARD TEMP ~ oC

ESSS/4−230 GALLON TANK CONFIGURATION
STANDARD DAY ZERO HEADWIND
HIRSS SUPPRESSED MODE
FUEL LOAD = 8,300 LB (JP4)
GROSS WEIGHT = 22,000 LB
80 LB WARM UP
BLEED AIR OFF

15
0

APPROX FLIGHT TIME ~ HRS

12000

11000

10000

FUEL REMAINING ~ LBS

9000
APPROX FULL FUEL
8000

7000
ADEQUATE
RANGE

6000

FU
EL

5000

4000

INADEQUATE RANGE
ABORT MISSION

3000

RE
M
AI
NI
NG

2000
X

INADEQUATE
RANGE

1000
LOW FUEL LIGHTS

DATA BASIS:

0
FLIGHT TEST

200

400

600

800

1000

DISTANCE TRAVELED ~ NM

Figure 7-39. Assault Mission Profile (4 - 230 Gallon Tanks)
7-150

1200

AA0610A
SA

TM 1-1520-237-10

ASSAULT MISSION PROFILE FOR MAX RANGE

ESSS/2−230 GALLON TANK CONFIGURATION
STANDARD DAY ZERO HEADWIND
HIRSS SUPPRESSED MODE
FUEL LOAD = 5,300 LB (JP4)
GROSS WEIGHT = 22,000 LB
80 LB WARM UP
BLEED AIR OFF

(RECOMMENDED AIRSPEED)
108 KIAS

102 KIAS

100 KIAS

95 KIAS
(~65%)

(~58%)
−1

PRESS ALT

8
(~69%)

6
(~70%)
4

(APPROX TRQ ~ %)
(~76%)

STANDARD TEMP ~ oC

−5

0
0

APPROX FLIGHT TIME ~ HRS
6000

5500

5000

4500
FU

FUEL REMAINING ~ LBS

4000

ADEQUATE
RANGE

3500

3000
INADEQUATE RANGE
ABORT MISSION
2500

2000

1500

AX
M

ON
SI
IS
M

US
DI
RA

INADEQUATE
RANGE

1000

500
LOW FUEL LIGHTS
0
0

DATA BASIS:

100
FLIGHT TEST

200

300

400

500

600

DISTANCE ~ NM

700

AA0609A
SA

Figure 7-40. Assault Mission Profile (2 - 230 Gallon Tanks)
7-151/(7-152 Blank)

TM 1-1520-237-10

CHAPTER 7A
PERFORMANCE DATA

701C

Section I INTRODUCTION
7A.1 PURPOSE.
NOTE
Chapter 7A contains performance data for
aircraft equipped with T700-GE-701C engines. Performance data for other models are
contained in Chapter 7. Users are authorized
to remove whichever chapter is not applicable to their model aircraft, and are not required to carry both chapters on board.

Section
and
Figure
Number

Title

Page

INTRODUCTION ...................

7A-1

Temperature Conversion
Chart.........................................

7A-5

MAXIMUM TORQUE
AVAILABLE...........................

7A-6

Aircraft Torque Factor
(ATF) .......................................

7A-7

Torque Conversion
Chart.........................................

7A-9

Maximum Torque
Available ..................................

7A-10

7A-5

Dual Engine Torque
Limit.........................................

7A-12

(2) Situations requiring maximum performance will be
more readily recognized.

III

HOVER....................................

7A-13

7A-6

Hover - Clean .........................

7A-14

(3) Familiarity with the data will allow performance to
be computed more easily and quickly.

7A-7

Hover - High Drag ..................

7A-16

CRUISE ...................................

7A-17

(4) Experience will be gained in accurately estimating
the effects of variables for which data are not presented.

7A-8

Sample Cruise Chart................

7A-19

7A-9

Cruise - Pressure Altitude Sea Level .................................

7A-20

Cruise High Drag-Pressure
Altitude - Sea Level ................

7A-26

Cruise - Pressure Altitude 2,000 Feet ................................

7A-32

Cruise High Drag-Pressure
Altitude - 2,000 Feet ...............

7A-38

Cruise - Pressure Altitude 4,000 Feet ................................

7A-44

Cruise High Drag-Pressure
Altitude - 4,000 Feet ...............

7A-50

a. The purpose of this chapter is to provide the best
available performance data for the UH-60L. Regular use of
this information will enable you to receive maximum safe
utilization of the helicopter. Although maximum performance is not always required, regular use of this chapter is
recommended for these reasons:
(1) Knowledge of your performance margin will allow
you to make better decisions when unexpected conditions
or alternate missions are encountered.

b. The information is primarily intended for mission
planning and is most useful when planning operations in
unfamiliar areas or at extreme conditions. The data may
also be used in flight, to establish unit or area standard
operating procedures, and to inform ground commanders of
performance/risk tradeoffs.
7A.2 CHAPTER 7A INDEX.
The following index contains a list of the sections, titles,
figure numbers, subjects and page numbers of each performance data chart contained in this chapter.

II
7A-2
7A-3
7A-4

7A-10
7A-11
7A-12
7A-13
7A-14

7A-1

TM 1-1520-237-10

Section
and
Figure
Number
7A-15
7A-16
7A-17
7A-18
7A-19
7A-20
7A-21
7A-22

Title

Page

Section
and
Figure
Number

Title

Page

Cruise - Pressure Altitude 6,000 Feet ................................

7A-33

External Load Drag ................. 7A-138

7A-56

7A-34

Cruise High Drag-Pressure
Altitude - 6,000 Feet ...............

Typical High
Drag Configurations ................ 7A-139

7A-62

VII

CLIMB - DESCENT............... 7A-140

Cruise - Pressure Altitude 8,000 Feet ................................

7A-68

7A-35

Climb/Descent.......................... 7A-141

Cruise High Drag-Pressure
Altitude - 8,000 Feet ...............

7A-36

7A-74

Climb/Descent High Drag ................................ 7A-142

Cruise - Pressure Altitude 10,000 Feet ..............................

VIII

FUEL FLOW........................... 7A-143

7A-80

7A-37

Cruise High Drag-Pressure
Altitude - 10,000 Feet .............

Single/Dual Engine
Fuel Flow................................. 7A-144

7A-85

Cruise - Pressure Altitude 12,000 Feet ..............................

AIRSPEED SYSTEM
CHARACTERISTICS ............. 7A-145

7A-90

7A-38

Airspeed Correction Chart ...... 7A-146

7A-39

Airspeed Correction
Chart - High Drag ................... 7A-147

Cruise High Drag-Pressure
Altitude - 12,000 Feet .............

7A-95

7A-23

Cruise - Pressure Altitude 14,000 Feet .............................. 7A-100

SPECIAL MISSION
PERFORMANCE .................... 7A-148

7A-24

Cruise High Drag-Pressure
Altitude - 14,000 Feet ............ 7A-105

7A-40

Self Deployment Mission
Profile....................................... 7A-150

7A-25

Cruise - Pressure Altitude 16,000 Feet .............................. 7A-110

7A-41

Assault Mission Profile
(4 - 230 Gallon Tanks)............ 7A-152

7A-26

Cruise High Drag-Pressure
Altitude - 16,000 Feet ............. 7A-114

7A-42

Assault Mission Profile
(2 - 230 Gallon Tanks)............ 7A-153

7A-27

Cruise - Pressure Altitude 18,000 Feet .............................. 7A-118

7A-28

Cruise High Drag-Pressure
Altitude - 18,000 Feet ............. 7A-122

7A-29

Cruise - Pressure Altitude 20,000 Feet .............................. 7A-126

7A-30

Cruise High Drag-Pressure
Altitude - 20,000 Feet ............. 7A-130

OPTIMUM CRUISE ............... 7A-134

7A-31

Optimum Altitude For
Maximum Range ..................... 7A-135

7A-32

Optimum Altitude For
Maximum Range High Drag ................................ 7A-136

DRAG ...................................... 7A-137

7A-2

Change 2

7A.3 GENERAL.
The data presented covers the maximum range of conditions and performance that can reasonably be expected.
In each area of performance, the effects of altitude, temperature, gross weight, and other parameters relating to that
phase of flight are presented. In addition to the presented
data, your judgment and experience will be necessary to
accurately obtain performance under a given set of circumstances. The conditions for the data are listed under the title
of each chart. The effects of different conditions are discussed in the text accompanying each phase of performance. Where practical, data are presented at conservative
conditions. However, NO GENERAL CONSERVATISM
HAS BEEN APPLIED. All performance data presented are
within the applicable limits of the helicopter. All flight per

TM 1-1520-237-10

formance data are based on JP-4 fuel. The change in fuel
flow and torque available, when using JP-5 or JP-8 aviation
fuel or any other approved fuels is insignificant.
7A.4 LIMITS.

7A.7 PERFORMANCE DATA BASIS - CLEAN.
The data presented in the performance charts are primarily derived for a clean UH-60L aircraft and are based on
U. S. Army test data. The clean configuration assumes all
doors and windows are closed and includes the following
external configuration:

CAUTION

Exceeding operating limits can cause permanent damage to critical components.
Overlimit operation can decrease performance, cause early failure, or failure on a
subsequent flight.
Applicable limits are shown on the charts. Performance
generally deteriorates rapidly beyond limits. If limits are
exceeded, minimize the amount and time. Enter the maximum value and time above limits on DA Form 2408-13-1
so proper maintenance action can be taken.
7A.5 USE OF CHARTS.
7A.5.1 Data Basis. The type of data used is indicated at
the bottom of each performance chart under DATA BASIS.
The data provided generally is based on one of three categories:
a. Flight test data. Data obtained by flight test of the
helicopter by experienced flight test personnel at precise
conditions using sensitive calibrated instruments.
b. Calculated data. Data based on tests, but not on flight
test of the complete helicopter.
c. Estimated data. Data based on estimates using aerodynamic theory or other means but not verified by flight
test.
7A.5.2 Specific Conditions. The data presented is accurate only for specific conditions listed under the title of
each chart. Variables for which data is not presented, but
which may affect that phase of performance, are discussed
in the text. Where data is available or reasonable estimates
can be made, the amount that each variable affects performance will be given.
7A.6 PERFORMANCE DISCREPANCIES.
Regular use of this chapter will allow you to monitor
instrument and other helicopter systems for malfunction, by
comparing actual performance with planned performance.
Knowledge will also be gained concerning the effects of
variables for which data is not provided, thereby increasing
the accuracy of performance predictions.

a. Fixed provisions for the External Stores Support System (ESSS).
b. Main and tail rotor deice system.
c. Mounting brackets for IR jammer and chaff dispenser.
d. The Hover Infrared Suppressor System (HIRSS) with
baffles installed.
e. Includes wire strike protection system.
NOTE
Aircraft which have an external configuration which differs from the clean configuration may be corrected for drag differences
on cruise performance as discussed in Section VI DRAG.
7A.8 PERFORMANCE DATA BASIS - HIGH DRAG.
The data presented in the high drag performance charts
are primarily derived for the UH-60L with the ESSS system installed and two 230-gallon tanks mounted on the outboard pylons, and are based on U. S. Army test data. The
high drag configuration assumes all doors and windows are
closed and includes the following external configuration:
a. External stores support system installed.
b. Two 230–gallon tanks mounted on the outboard pylons.
c. Inboard vertical pylons empty.
d. IR jammer and chaff dispenser installed.
e. Hover Infrared Suppressor System (HIRSS) with
baffles are installed.
f. Main and tail rotor deice and wire strike protection
systems are installed.

Change 10

7A-3

TM 1-1520-237-10

NOTE
Aircraft with an external configuration that
differs from the high drag configuration
baseline may be corrected for differences in
cruise performance as discussed in Section
VI DRAG.

based on flight test data obtained with the complete volcano
system installed, to include all of the canisters and mines.
The drag correction factor may be used to provide a conservative estimate of cruise performance for volcano configurations which do not include all of the canisters and
mines.
7A.9 FREE AIR TEMPERATURES.

g. VOL Use the high drag configuration hover charts
to determine hover performance with the volcano system
installed. Use the high drag cruise charts and the volcano
drag correction factor to determine cruise performance with
volcano installed. The volcano drag correction factor is

7A-4

A temperature conversion chart (Figure 7A-1) is included for the purpose of converting Fahrenheit temperature to Celsius.

TM 1-1520-237-10

TEMPERATURE CONVERSION
EXAMPLE
WANTED:
FREE AIR TEMPERATURE IN DEGREES CELSIUS

KNOWN:
FREE AIR TEMPERATURE = 32oF

METHOD:
ENTER FREE AIR TEMPERATURE HERE
MOVE RIGHT TO DIAGONAL LINE
MOVE DOWN TO DEGREES CELSIUS SCALE
READ FREE AIR TEMPERATURE = 0oC

140

120

100

FAT ~ oF

−20

−40

−60

−80
−60

−50

−40

−30

−20

−10

FAT ~ oC
AA0674
SA

Figure 7A-1. Temperature Conversion Chart

7A-5

TM 1-1520-237-10

Section II MAXIMUM TORQUE AVAILABLE
7A.10 TORQUE FACTOR METHOD.
The torque factor method provides an accurate indication of available power by incorporating ambient temperature effects on degraded engine performance. This section
presents the procedure to determine the maximum dual- or
single-engine torque available. Specification power is defined for a newly delivered low time engine. The aircraft
HIT log forms for each engine provide the engine and aircraft torque factors which are obtained from the maximum
power check and recorded to be used in calculating maximum torque available.
7A.10.1 Torque Factor Terms. The following terms
are used when determining the maximum torque available
for an individual aircraft:
a. Torque Ratio (TR). The ratio of torque available to
specification torque at the desired ambient temperature.
b. Engine Torque Factor (ETF). The ratio of an individual engine torque available to specification torque at reference temperature of 35°C (95°F). The ETF is allowed to
range from .85 to 1.0.
c. Aircraft Torque Factor (ATF). The ratio of an individual aircraft’s power available to specification power at a
reference temperature of 35°C (95°F). The ATF is the average of the ETF’s of both engines and its value is allowed
to range from 0.9 to 1.0.
7A.10.2 Torque Factor Procedure. The use of the
ATF or ETF to obtain the TR from Figure 7A-2 for ambient temperatures between -15°C (5°F) and 35°C (95°F) is
shown by the example. The ATF and ETF values for an
individual aircraft are found on the engine HIT Log. The
TR always equals 1.0 for ambient temperatures of -15°C
(5°F) and below, and the TR equals the ATF or ETF for
temperatures of 35°C (95°F) and above.
7A.11 TORQUE AVAILABLE.
a. This section presents the maximum dual-engine
torque available for the 2.5-minute, 10-minute and 30minute limits at zero airspeed and 100% RPM R for the
operational range of pressure altitude and FAT. The singleand dual-engine transmission limits for continuous operation are also shown and should not be exceeded.

7A-6

Change 6

CAUTION

Do not exceed the UH-60L DUAL ENGINE TORQUE LIMITS in Chapter 5.
These torque limits are presented in Figure 7A-5 and on the TORQUE PLACARD mounted on the instrument panel.
b. When the TR equals 1.0, the torque available may be
read directly from the torque available per engine scales.
When the TR is less than 1.0, the actual torque available is
determined by multiplying the specification torque available by the TR (example for TR = 0.98: 90% TRQ x 0.98
= 88.2% TRQ). The torque conversion chart (Figure 7A-3)
is provided to convert specification data to actual torque
available.
7A.11.1 Torque Available - 2.5 Minutes. Figure 7A-4
presents the specification torque available at 903°C TGT
per engine for the 2.5 minute limit. Contingency (2.5
Minute) Power is automatically available when any one
engine torque is less than 50% or when the pilot selects
DEC LOCKOUT and manually maintains the 2.5 minute
TGT limit.
7A.11.2 Torque Available - 10 Minutes. Figure 7A-4
presents the specification torque available per engine for
the 10 minute limit. This is the maximum dual-engine
torque available and is set by the TGT limiter in dualengine operation. For one engine operation, the pilot must
maintain the 10 minute TGT limit.
7A.11.3 Torque Available - 30 Minutes. Figure 7A-4
presents the specification torque available per engine for
the 30 minute limit. The pilot must manually maintain the
30 minute TGT limit.
7A.12 ENGINE BLEED AIR.
With engine bleed air on, the available torque per engine
is reduced as follows:
a. Engine anti-ice on - 18% TRQ
(example: 90% TRQ - 18% TRQ = 72% TRQ).
b. Cockpit heater on - 4% TRQ.

TM 1-1520-237-10

c. Both on - 22% TRQ.

7A.13 INFRARED SUPPRESSOR SYSTEM.
When the IR suppressor is OPERATING IN THE BENIGN MODE (exhaust baffles removed) the torque available is increased about 1% TRQ.

Change 6

7A-6.1/(7A-6.2 Blank)

TM 1-1520-237-10

TORQUE FACTOR
T700−GE−701C ENGINE 100% RPM R
TORQUE FACTOR ~ ATF OR ETF
40

.85

.86

.87

.88

.89

.90

.91

.92

.93

.94

.95

.96

.97

.98

.99

1.0

FOR FAT’S
OF 35oC AND
ABOVE:
TR = ATF

35
2

FREE AIR TEMPERATURE ~ o C

−5

−10

FOR FAT’S
OF −15oC AND
BELOW:
TR = 1.0

−15

−20
.85

.86

.87

.88

.89

.90

.91

.92

.93

.94

.95

.96

.97

.98

.99

1.0

.954 3

TORQUE RATIO ~ TR

EXAMPLE
WANTED:

METHOD:

TORQUE RATIO AND MAXIMUM TORQUE AVAILABLE −
10−MINUTE LIMIT

TO OBTAIN TORQUE RATIO:

KNOWN:
ATF = .95
PRESSURE ALTITUDE = 6000 FT.
FAT = 30oC

1. ENTER TORQUE FACTOR CHART AT KNOWN FAT
2. MOVE RIGHT TO THE ATF VALUE
3. MOVE DOWN, READ TORQUE RATIO = .954.
TO DETERMINE SPECIFICATION TORQUE AVAILABLE −
10−MINUTE LIMITS:
4. ENTER MAXIMUM TORQUE AVAILABLE CHART AT
KNOWN FAT (FIGURE 7A−4).
5. MOVE RIGHT TO KNOWN PRESSURE ALTITUDE
6. MOVE DOWN, READ SPECIFICATION TORQUE = 98%.
TO OBTAIN ACTUAL TORQUE VALUE AVAILABLE FROM THE
TORQUE CONVERSION CHART:

NOTE
EITHER OF THE TWO TORQUE
AVAILABLE CHARTS MAY BE USED.
MAXIMUM ALLOWABLE DUAL ENGINE
TORQUE LIMITS SHALL NOT BE
EXCEEDED.

DATA BASIS:

CALCULATED

7. ENTER TORQUE CONVERSION CHART FIGURE 7A−3 AT %
TORQUE OBTAINED FROM FIGURE 7A−4.
8. MOVE UP TO TORQUE RATIO OBTAINED FROM FIGURE 7A−2
9. MOVE LEFT, READ MAXIMUM TORQUE AVAILABLE − 10
MINUTE LIMIT = 93%.
10. ENTER DUAL−ENGINE TORQUE LIMIT CHART FIGURE 7A−5
AT 30oC. MOVE RIGHT TO INTERSECTION AT 6,000 FT. PA.
AA0692C
MOVE DOWN TO READ 90.4% TORQUE.
SA

Figure 7A-2. Aircraft Torque Factor (ATF)

7A-7

TM 1-1520-237-10

7A.14 DUAL-ENGINE TORQUE LIMITS.
Helicopters prior to S/N 91–26354 that are not equipped
with improved main rotor flight controls are further restricted above 80 KIAS to dual-engine continuous torque

7A-8

Change 6

limits as indicated by a placard on the instrument panel.
Figure 7A-5 graphically presents the dual-engine torque
limits for use with the torque available charts.

TM 1-1520-237-10

TORQUE CONVERSION
TORQUE RATIO
135

1.00
.98

130
.96
.94
125
.92
.90

120

.88
.86
115
.84

ACTUAL TORQUE AVAILABLE ~ %

110

105

100

85
8
80

50
7
45
60

100

110

120

130

TORQUE AVAILABLE PER ENGINE (SPECIFICATION TORQUE) ~ %
AA1636A
SA

Figure 7A-3. Torque Conversion Chart

Change 6

7A-9

TM 1-1520-237-10

MAXIMUM TORQUE AVAILABLE − 2.5−MINUTE LIMIT
T700−GE−701C
HIRSS (BAFFLES INSTALLED)
100% RPM R
BLEED AIR OFF
ZERO AIRSPEED
2−ENGINE TORQUE
LIMIT ABOVE 80 KIAS

1−ENGINE
TRANSMISSION
LIMIT

2−ENGINE
TRANSMISSION
LIMIT

60
0
2
50

PRESSURE ALTITUDE
~ 1000 FT

FREE AIR TEMPERATURE (FAT) ~ OC

10
30

ENGINE
HIGH AMBIENT
TEMPERATURE LIMIT

12
14

16
18

−10

−20

−30

−40

ENGINE
LOW AMBIENT
TEMPERATURE
LIMIT

−50
20

−60

DATA BASIS:

ENGINE MANUFACTURER
SPEC.

100

110

120

TORQUE AVAILABLE PER ENGINE ~ %

Figure 7A-4. Maximum Torque Available (Sheet 1 of 3)

7A-10

Change 6

130
AA1000_4A
SAF

TM 1-1520-237-10

MAXIMUM TORQUE AVAILABLE − 10−MINUTE LIMIT
T700−GE−701C
HIRSS (BAFFLES INSTALLED)
100% RPM R
BLEED AIR OFF
ZERO AIRSPEED
2−ENGINE TORQUE
LIMIT ABOVE 80 KIAS
60

0
2
4

PRESSURE ALTITUDE
~ 1000 FT

ENGINE
HIGH AMBIENT TEMPERATURE
LIMIT

6
8

10
12

FREE AIR TEMPERATURE (FAT) ~ OC

1−ENGINE
TRANSMISSION
LIMIT

2−ENGINE
TRANSMISSION
LIMIT

4 30

16
18

−10

−20

−30

−40

−50
20

−60

DATA BASIS:

ENGINE MANUFACTURER
SPEC.

6
98

100

4
110

2
120

0
130

6
AA1000_1A

TORQUE AVAILABLE PER ENGINE ~ %

Figure 7A-4. Maximum Torque Available (Sheet 2 of 3)

Change 6

7A-11

TM 1-1520-237-10

MAXIMUM TORQUE AVAILABLE − 30−MINUTE LIMIT
T700−GE−701C
HIRSS (BAFFLES INSTALLED)
100% RPM R
BLEED AIR OFF
ZERO AIRSPEED
2−ENGINE TORQUE
LIMIT ABOVE 80 KIAS
60

1−ENGINE
TRANSMISSION
LIMIT

2−ENGINE
TRANSMISSION
LIMIT

0
4

PRESSURE ALTITUDE
~1000 FT
40

8
10

FREE AIR TEMPERATURE (FAT) ~ OC

12
30

14
16

20
18
10

−10

−20

−30

−40

−50
20

−60
60

DATA BASIS:
ENGINE MANUFACTURER SPEC.

100

110

TORQUE AVAILABLE PER ENGINE ~ %

Figure 7A-4. Maximum Torque Available (Sheet 3 of 3)

7A-12

Change 6

120

130
AA1000_2A
SA

TM 1-1520-237-10

DUAL−ENGINE TORQUE LIMITS ABOVE 80 KIAS
T−700−GE−701C 100% RPM R
FOR AIRCRAFT WITH TORQUE PLACARD ONLY
60
PRESSURE ALTITUDE ~ 1000 FT
20

FREE AIR TEMPERATURE ~ oC

−10

−20

−30

−40

−50

−60
40

DATA BASIS:

FLIGHT TEST

100

MAXIMUM ALLOWABLE DUAL−ENGINE TORQUE ~ %

AA1255A
SA

Figure 7A-5. Dual-Engine Torque Limit

Change 6

7A-12.1/(7A-12.2 Blank)

TM 1-1520-237-10

Section III HOVER
7A.15 HOVER CHART.
NOTE
VOL For performance calculations with
the volcano system installed, use the applicable high drag performance charts.

a. The primary use of the chart (Figures 7A-6 and 7A-7)
is illustrated by part A of the example. To determine the
torque required to hover, it is necessary to know pressure
altitude, free air temperature, gross weight, and desired
wheel height. Enter the upper right grid at the known free
air temperature, move right to the pressure altitude, move
down to gross weight. For OGE hover, move left to torque
per engine scale and read torque required. For IGE hover,
move left to desired wheel height, deflect down and read
torque required for dual-engine or single-engine operation.
The IGE wheel height lines represent a compromise for all
possible gross weights and altitude conditions. A small
torque error up to 6 3% torque may occur at extreme temperature and high altitude. This error is more evident at
lower wheel heights.
b. In addition to the primary use, the hover chart (Figure
7A-6) may be used to predict maximum hover height. To
determine maximum hover height, it is necessary to know
pressure altitude, free air temperature, gross weight, and
maximum torque available. Enter the known free air temperature move right to the pressure altitude, move down to

gross weight, move left to intersection with maximum
torque available and read wheel height. This wheel height
is the maximum hover height.
c. The hover chart may also be used to determine maximum gross weight for hover at a given wheel height, pressure altitude, and temperature as illustrated in method B of
the example (Figure 7A-6). Enter at known free air temperature, move right to the pressure altitude, then move
down and establish a vertical line on the lower grid. Now
enter lower left grid at maximum torque available. Move up
to wheel height, then move right to intersect vertical line
from pressure altitude/FAT intersection. Interpolate from
gross weight lines to read maximum gross weight at which
the helicopter will hover.
7A.16 EFFECTS OF BLADE EROSION KIT.
With the blade erosion kit installed, it will be necessary
to make the following corrections. Multiply the torque required to hover determined from the charts by 1.02. (Example: If indicated torque is 90%, multiply 90 x 1.02 =
91.8% actual torque required.) Multiply the maximum gross
weight to hover obtained from the charts by 0.98. (Example: If gross weight is 22,000 lb, multiply by 0.98 =
21,560 lb actual gross weight to hover.) When determining
maximum hover wheel height, enter the chart at 1.02 x
gross weight. (Example: If gross weight is 20,000 lb, multiply 20,000 x 1.02 = 20,400 lb).

7A-13

TM 1-1520-237-10

EXAMPLE A
WANTED:
TORQUE REQUIRED TO HOVER OGE AND AT A 10-FOOT WHEEL HEIGHT
KNOWN:
FAT = 30°C
PRESSURE ALTITUDE = 3,000 FEET
GROSS WEIGHT = 19,500 POUNDS
METHOD:
ENTER HOVER CHART AT KNOWN FAT. MOVE RIGHT TO PRESSURE ALTITUDE, MOVE DOWN
THROUGH GROSS WEIGHT LINES TO DESIRED GROSS WEIGHT. MOVE LEFT TO INDICATE
TORQUE/ENGINE % (OGE) SCALE AND READ OGE HOVER TORQUE (95%). MOVE DOWN
FROM INTERSECTION OF 10-FOOT HOVER LINE AND HORIZONTAL LINE TO READ TORQUE
REQUIRED TO HOVER 10 FEET (80%).

EXAMPLE B
WANTED:
MAXIMUM GROSS WEIGHT TO HOVER OGE
KNOWN:
ATF = 1.0
FAT = 20°C
PRESSURE ALTITUDE = 5,000 FEET
MAXIMUM TORQUE AVAILABLE = 107%
METHOD:
ENTER INDICATED TORQUE/ENGINE (IGE) SCALE AT MAXIMUM TORQUE AVAILABLE
(107%), MOVE UP TO OGE LINE. ENTER CHART AT KNOWN FAT (20°C). MOVE RIGHT TO
PRESSURE ALTITUDE LINE. MOVE DOWN FROM PRESSURE ALTITUDE LINE AND MOVE
RIGHT FROM OGE LINE. WHERE LINES INTERSECT, READ MAXIMUM GROSS WEIGHT TO
HOVER OGE (20,500 lb).

Figure 7A-6. Hover - Clean (Sheet 1 of 2)

7A-14

TM 1-1520-237-10

HOVER

HOVER
CLEAN
T701C(2)

CLEAN CONFIGURATION 100% RPM R
ZERO WIND

PRESSURE ALTITUDE ~ 1000 FT

5
WHEEL
HEIGHT ~ FT

130

16
18

20
B
20

0
−20
−40

125
120

−60
20

FREE AIR TEMP ~ O C

NOTE
FOR LOW WIND CONDITIONS
AIRCRAFT SHOULD BE HEADED
INTO WIND. 3−5 KT CROSSWIND
OR TAILWIND MAY INCREASE
TORQUE REQUIRED BY UP TO
4% OVER ZERO WIND VALUES

−2

23.5 23

40
DUAL ENGINE TRANS LIMIT

OGE
16

115

105
15
100
95
SINGLE ENGINE
TRANS LIMIT

14
85

DUAL ENGINE TRANS LIMIT

TORQUE PER ENGINE ~ % (OGE)

110

80
75
70
65

60
GW ~
1000 LB

55
B
50
40

100

110

120

TORQUE PER ENGINE ~ % (IGE)
80

DATA BASIS :

100

120

140

SINGLE ENGINE TORQUE ~ %
AA2146D

FLIGHT TEST

SAF

Figure 7A-6. Hover - Clean (Sheet 2 of 2)

Change 7

7A-15

TM 1-1520-237-10

HOVER

HOVER
ESSS
T701C (2)

HIGH DRAG CONFIGURATION 100% RPM R
ZERO WIND
60

FOR LOW WIND CONDITIONS
AIRCRAFT SHOULD BE HEADED
INTO WIND. 3−5 KT CROSSWIND
OR TAILWIND MAY INCREASE
TORQUE REQUIRED BY UP TO
4% OVER ZERO WIND VALUES
WHEEL
HEIGHT ~ FT

135

PRESSURE ALTITUDE ~ 1000 FT
0
2
4
6
8
10

FREE AIR TEMP ~ O C

NOTE

−2

20
20

0
−20
−40
−60

24.5 24

130
40

125
120

DUAL ENGINE TRANS LIMIT

OGE

110

15
105
SINGLE ENGINE
TRANS LIMIT

100
95
14

DUAL ENGINE TRANS LIMIT

TORQUE PER ENGINE ~ % (OGE)

115

85
80
75
70

GW ~
1000 LB

65
60
55
40

100

110

120

TORQUE PER ENGINE ~ % (IGE)
80

100

DATA BASIS :

120

140

SINGLE ENGINE TORQUE ~ %
AA2147C

FLIGHT TEST

SAF

Figure 7A-7. Hover - High Drag

7A-16

Change 7

TM 1-1520-237-10

Section IV CRUISE
7A.17 DESCRIPTION.
The cruise charts (Figures 7A-8 through 7A-30) present
torque required and total fuel flow as a function of airspeed,
altitude, temperature, and gross weight at 100% rotor speed.
Scales for both true airspeed and indicated airspeed are
presented. The baseline aircraft configurations for these
charts are 9clean and high drag 9 configuration as defined in
Section I. Each cruise chart also presents the change in
torque ( TRQ) required for 10 sq. ft. of additional flat
plate drag with a dashed line on a separate scale. This line
is utilized to correct torque required for external loads as
discussed in Section VI DRAG. Maximum level flight airspeed (Vh) is obtained at the intersection of gross weight
arc and torque available - 30 minutes or the transmission
torque limit, whichever is lower. Airspeeds that will produce maximum range, maximum endurance, and maximum
rate of climb are also shown. Cruise charts are provided
from sea level to 20,000 feet pressure altitude in units of
2,000 feet. Each figure number represents a different altitude. The charts provide cruise data for free air temperatures from -50° to +60°C, in units of 10°. Charts with
FAT’s that exceed the engine ambient temperature limits
by more than 10°C are deleted.
7A.18 USE OF CHARTS.
The primary uses of the charts are illustrated by the
examples of Figure 7A-8. To use the charts, it is usually
necessary to know the planned pressure altitude, estimated
free air temperature, planned cruise speed, TAS, and gross
weight. First, select the proper chart on the basis of pressure altitude and FAT. Enter the chart at the cruise airspeed, IAS, move horizontal and read TAS, move horizontal to the gross weight, move down and read torque
required, and then move up and read associated fuel flow.
Maximum performance conditions are determined by entering the chart where the maximum range line or the maximum endurance and rate of climb line intersects the gross
weight line; then read airspeed, fuel flow, and torque required. Normally, sufficient accuracy can be obtained by
selecting the chart nearest the planned cruising altitude and
FAT or, more conservatively, by selecting the chart with
the next higher altitude and FAT. If greater accuracy is
required, interpolation between altitudes and/or temperatures is permissible. To be conservative, use the gross
weight at the beginning of the cruise flight. For greater
accuracy on long flights, however, it is preferable to determine cruise information for several flight segments to allow
for the decreasing gross weight.

a. Airspeed. True and indicated airspeeds are presented
at opposite sides of each chart. On any chart, indicated
airspeed can be directly converted to true airspeed (or vice
versa) by reading directly across the chart without regard
for the other chart information.
b. Torque. Since pressure altitude and temperature are
fixed for each chart, torque required varies according to
gross weight and airspeed. The torque and torque limits
shown on these charts are for dual-engine operation. The
maximum torque available is presented on each chart as
either the transmission torque limit or torque available - 30
minute for both ATF-1.0 and 0.9 values. The maximum
torque available for aircraft with an ATF value between
these must be interpolated. The continuous torque available
values shown represent the minimum torque available for
ATFs of 1.0. For ATFs less than 1.0, maximum continuous
torque available may be reduced. The dual-engine torque
limit placard value is presented below the torque scale of
each chart when applicable. An increase or decrease in
torque required because of a drag area change is calculated
by adding or subtracting the change in torque from the
torque on the curve, and then reading the new fuel flow
total.
c. Fuel Flow. Fuel flow scales are provided opposite the
torque scales. On any chart, torque may be converted directly to fuel flow without regard to other chart information. Data shown in this section is for two-engine operation.
For one-engine fuel flow, refer to Section VIII FUEL
FLOW.
(1) With bleed-air extracted, fuel flow increases:
(a) Engine anti-ice on - About 100 lbs/hr. Example:
(760 lbs/hr + 100 lbs/hr = 860 lbs/hr.)
(b) Heater on - About 12 lbs/hr.
(c) Both on - About 112 lbs/hr.
(2) When the hover IR suppressor system is operating
in the benign mode (exhaust baffles removed), the dualengine fuel flow will decrease about 14 lbs/hr.
d. Maximum Range. The maximum range lines (MAX
RANGE) indicate the combinations of gross weight and
airspeed that will produce the greatest flight range per
pound of fuel under zero wind conditions. When maximum

Change 10

7A-17

TM 1-1520-237-10

range airspeed line is above the maximum torque available,
the resulting maximum airspeed should be used for maximum range. A method of estimating maximum range speed
in winds is to increase IAS by 2.5 knots per each 10 knots
of effective headwind (which reduces flight time and minimizes loss in range) and decrease IAS by 2.5 knots per 10
knots of effective tailwind for economy.
e. Maximum Endurance and Rate of Climb. The maximum endurance and rate of climb lines (MAX END and
R/C) indicate the combinations of gross weight and airspeed that will produce the maximum endurance and the
maximum rate of climb. The torque required for level flight
at this condition is a minimum, providing a minimum fuel
flow (maximum endurance) and a maximum torque change
available for climb (maximum rate of climb).
f. Change in Frontal Area. Since the cruise information
is given for the 9clean and high drag configuration,9 adjustments to torque should be made when operating with external sling loads or aircraft external configuration changes.
To determine the change in torque, first obtain the appropriate multiplying factor from the drag load chart (Figure
7A-33 or Table 7A-1), then enter the cruise chart at the
planned cruise speed TAS, move right to the broken
TRQ line, and move up and read TRQ. Multiply
TRQ by the multiplying factor to obtain change in
torque, then add or subtract change in torque from torque
required for the primary mission configuration. Enter the
cruise chart at resulting torque required, move up, and read
fuel flow. If the resulting torque required exceeds the governing torque limit, the torque required must be reduced to
the limit. The resulting reduction in airspeed may be found
by subtracting the change in torque from the limit torque;
then enter the cruise chart at the reduced torque, and move
up to the gross weight. Move left or right to read TAS or
IAS. The engine torque setting for maximum range obtained from the clean configuration cruise chart will generally result in cruise at best range airspeed for the higher
drag configuration. To determine the approximate airspeed
for maximum range for alternative or external load configurations, reduce the value from the cruise chart by 6 knots
F. For
for each 10 square foot increase in drag area,
example, if both cabin doors are open the F increases 6
ft2 and the maximum range airspeed would be reduced by

7A-18

approximately 4 knots (6 Kts/10 ft2x6 ft2 = 3.6 Kts). Only
the high drag cruise charts have data for gross weights
above 22,000 pounds. For external cargo hook load operations in excess of 8,000 pounds that attain gross weights
from 22,000 to 23,500 pounds it will be necessary to use
the high drag cruise charts. If the external stores support
system (ESSS) and the two 230-gallon tanks are not installed and the estimated drag value for the cargo hook load
is greater than 14 square feet , it will be necessary to subtract 14 square feet of drag from the cruise chart drag value
when determining cruise performance. If the ESSS and the
two 230-gallon tanks are not installed and the drag is estimated to be 14 square feet or less, the high drag charts
should be used with no other correction.
g. Additional Uses. The low speed end of the cruise
chart (below 40 knots) is shown primarily to familiarize
you with the low speed power requirements of the helicopter. It shows the power margin available for climb or acceleration during maneuvers, such as NOE flight. At zero airspeed, the torque represents the torque required to hover
out of ground effect. In general, mission planning for low
speed flight should be based on hover out of ground effect.
7A.19 SINGLE-ENGINE.
a. The minimum or maximum single-engine speeds can
be determined by using a combination of the 701C torque
available and cruise charts. To calculate single-engine
speeds, first determine the torque available from Section II
at the TGT limit desired and divide by 2. (Example: 90%
TRQ 4 2 = 45% TRQ.)
b. Select the appropriate cruise chart for the desired
flight condition and enter the torque scale with the torque
value derived above. Move up to the intersection of torque
available and the mission gross weight arc, and read across
for minimum single-engine airspeed. Move up to the second intersection of torque and weight, and read across to
determine the maximum single-engine speed. If no intersections occur, there is no single-engine level flight capability for the conditions. Single-engine fuel flow at the desired 10 minute, 30 minute, continuous conditions may be
obtained by doubling the torque required from the cruise
chart and referring to Figure 7A-37.

TM 1-1520-237-10

CRUISE EXAMPLE
CLEAN CONFIGURATION
100% RPM R
FAT: 30O C ALT: 6,000 FT
TOTAL FUEL FLOW 100 LB/HR
6

WANTED

ATF=0.9

180

CONTINUOUS

EXAMPLE

ATF=1.0

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

160

170
A. CRUISE CONDITIONS FOR MAXIMUM RANGE
B. CONDITIONS FOR MAXIMUM ENDURANCE
C. MAXIMUM AIRSPEED IN LEVEL FLIGHT
D. DETERMINE TORQUE AND FUEL FLOW
REQUIRED TO CRUISE WITH CARGO
DOORS OPEN

150

160

140

150
130

KNOWN

130

METHOD
A. TURN TO CRUISE CHARTS NEAREST KNOWN
FLIGHT CONDITIONS, AT INTERSECTION
OF MAX RANGE LINE AND KNOWN VALUE OF
GROSS WEIGHT:
MOVE LEFT, READ TAS = 137 KTS
MOVE RIGHT, READ IAS = 121 KTS
MOVE DOWN, READ TORQUE = 63% TRQ
MOVE UP, READ TOTAL FUEL FLOW = 970 LBS / HR
B. AT INTERSECTION OF MAX END AND R / C
LINE AND KNOWN VALUE OF GROSS WEIGHT:
MOVE LEFT, READ TAS = 81 KTS
MOVE RIGHT, READ IAS = 66 KTS
MOVE DOWN, READ TORQUE = 40% TRQ
MOVE UP, READ TOTAL FUEL FLOW = 740 LBS / HR
C. THE 30−MINUTE TORQUE AVAILABLE FOR
AN ATF OF 0.95 IS ABOVE THE PLACARD
TORQUE LIMIT. THEREFORE, AT THE
INTERSECTION OF GROSS WEIGHT AND
90% TRQ:
MOVE LEFT, READ MAXIMUM TAS = 157 KTS
MOVE RIGHT, READ MAXIMUM IAS = 139 KTS
MOVE DOWN, READ MAXIMUM TORQUE = 90% TRQ
MOVE UP, READ TOTAL FUEL FLOW = 1250 LBS / HR
D. ENTER TRQ% PER 10 SQ FT SCALE AT 137 KTAS
MOVE UP READ TRQ = 9.0%
TURN TO DRAG TABLE IN SECTION VII
NOTE CARGO DOORS OPEN = 6.0 SQ FT F
AND HAS A DRAG MULTIPLYING FACTOR VALUE
OF 0.60, CALCULATE TOTAL TORQUE REQUIRED
USING THE CONDITIONS OF EXAMPLE A:
63% + (0.6 X 9.0%) = 68.4% TRQ
READ FUEL FLOW AT TOTAL TORQUE = 1025 LBS / HR

TRUE AIRSPEED ~ KTS

120

110

100

90
MAX END
AND R / C
80

120

110

100

INDICATED AIRSPEED ~ KTS

TRANSMISSION TORQUE LIMIT

140
FAT = 30OC
PRESSURE ALTITUDE = 6000 FT
GW = 17000 LBS
ATF = 0.95
PLACARD TORQUE LIMITS APPLY

TORQUE AVAILABLE ~ 30 MINUTES

MAX
RANGE

60
70
50
60
GW ~
1000 LB

50
12

40
20

30
20

10
10

0
20

TORQUE PER ENGINE ~ %

100
AA1266A
SA

Figure 7A-8. Sample Cruise Chart

7A-19

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

140
10

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

140

150

130

130
120

140

MAX
RANGE

120

130
110
110
120
100

100

80
MAX END
AND R / C
70

90
TRANSMISSION TORQUE LIMIT

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

90
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

100

110

40
12

14 16 18

GW ~
1000 LB

14 16 18

GW ~
1000 LB

30
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-9. Cruise - Pressure Altitude Sea Level (Sheet 1 of 6)
7A-20

AA1060_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

160

150

140
140

150
130

120

MAX
RANGE

110

130

120

110

100
100

80
MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

90
90
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

110

100

TRUE AIRSPEED ~ KTS

130

140

MAX
RANGE

50
40

40
30
GW ~
1000 LB

30
GW ~
1000 LB

20
12

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1060_2
SA

Figure 7A-9. Cruise - Pressure Altitude Sea Level (Sheet 2 of 6)
7A-21

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

160

150

140

140
140

MAX
RANGE

130

MAX
RANGE

130

130
120

120
120

110
110

100

100
90

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

50
40

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-9. Cruise - Pressure Altitude Sea Level (Sheet 3 of 6)
7A-22

AA1060_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
160

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

120

110

100

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

MAX END
AND R / C

130

100

MAX END
AND R / C

40
12

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

130

TRANSMISSION TORQUE LIMIT

MAX
RANGE

130

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1060_4
SA

Figure 7A-9. Cruise - Pressure Altitude Sea Level (Sheet 4 of 6)
7A-23

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

140

MAX
RANGE

130

130
120

120

100
100

MAX END
AND R / C
80

MAX END
AND R / C

GW ~
1000 LB
30

GW ~
1000 LB

100

TORQUE AVAILABLE
~30 MINUTES ATF = 0.9

110

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

~ CONTINUOUS

120

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

MAX
RANGE

40
30

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-9. Cruise - Pressure Altitude Sea Level (Sheet 5 of 6)
7A-24

AA1060_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 0 FT
60oC

50oC

IAS ~ KTS
14

TOTAL FUEL FLOW ~ 100 LB/HR
7

180

170

160

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
30

160
150

180

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

170
160

150

140
MAX
RANGE

100

ATF = 0.9

90
MAX END
AND R / C

TORQUE AVAILABLE ~ 30 MINUTES

50
12

GW ~
1000 LB

40
30

120

110

100

130

120

110

TRANSMISSION TORQUE LIMIT

110

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

120

TORQUE AVAILABLE ~ 30 MINUTES

130

140

130

~ CONTINUOUS

TRANSMISSION TORQUE LIMIT

140

MAX
RANGE

MAX END
AND R / C
70

100

TRUE AIRSPEED ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

50
12

GW ~
1000 LB

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1060_6
SA

Figure 7A-9. Cruise - Pressure Altitude Sea Level (Sheet 6 of 6)
7A-25

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

PRESS ALT: 0 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

160

150

140

140
130

140
130
130

120
MAX
RANGE

120

110

100
100

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

80
MAX END
AND R / C
70

TRUE AIRSPEED ~ KTS

110

100

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

120
MAX
RANGE

50
40

40
30

GW ~
1000 LB

24.5

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-10. Cruise High Drag - Pressure Altitude Sea Level (Sheet 1 of 6)
7A-26

AA1065_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

PRESS ALT: 0 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

160

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

160

150

150
140

140

140
130

130

130
120
MAX
RANGE

MAX
RANGE

120

110

100

100
90

80
MAX END
AND R / C
70

CONTINUOUS TORQUE LIMIT

90
CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

80
MAX END
AND R / C
70

TRUE AIRSPEED ~ KTS

170

50
40

30
12

14 16

24.5

GW ~
1000 LB

24.5

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1065_2
SA

Figure 7A-10. Cruise High Drag - Pressure Altitude Sea Level (Sheet 2 of 6)
7A-27

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

PRESS ALT: 0 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
6

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

180
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

170

160
160

160

150

150
140

140

140
130
MAX
RANGE

120

110

100

120

110

100

80
MAX END
AND R / C
70

90
80

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

130
MAX
RANGE

MAX END
AND R / C

50
40
12

24.5

GW ~
1000 LB

14 16

24.5

TRUE AIRSPEED ~ KTS

130

GW ~
1000 LB
20

30
20

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-10. Cruise High Drag - Pressure Altitude Sea Level (Sheet 3 of 6)
7A-28

AA1065_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

PRESS ALT: 0 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

180

160

170
10

160

170

150

160

140

150

150
140

140

130
130
MAX
RANGE

130

MAX
RANGE

120

110
100

100

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

80
MAX END
AND R / C

40
12

24.5

50
12

TRUE AIRSPEED ~ KTS

110

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

14 16

24.5
40

GW ~
1000 LB

GW ~
1000 LB
20

30
20

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1065_4
SA

Figure 7A-10. Cruise High Drag - Pressure Altitude Sea Level (Sheet 4 of 6)
7A-29

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

PRESS ALT: 0 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

180

160

180

170

170
150

160

150

140

130

150

120

100

110

90
CONTINUOUS TORQUE LIMIT

80
MAX END
AND R / C

MAX END
AND R / C

GW ~
1000 LB

24.5

100

TRUE AIRSPEED ~ KTS

100

110

CONTINUOUS TORQUE LIMIT

110

130

~ CONTINUOUS

120

TRUE AIRSPEED ~ KTS

MAX
RANGE

120

TORQUE AVAILABLE
~ 30 MINUTES ATF = 0.9

MAX
RANGE

130

140

30
20

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-10. Cruise High Drag - Pressure Altitude Sea Level (Sheet 5 of 6)
7A-30

AA1065_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
0 FT
T701C (2)

PRESS ALT: 0 FT
60oC

50oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

160
150

180

~ CONTINUOUS

170

150

140

170
160
150

130

MAX
RANGE

120

110

100

90
MAX END
AND R / C

110

100

90
CONTINUOUS TORQUE LIMIT

120

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

130

TORQUE AVAILABLE ~ 30 MINUTES

MAX
RANGE

80
MAX END
AND R / C

140
130

120

110

CONTINUOUS TORQUE LIMIT

140

100

TRUE AIRSPEED ~ KTS

ATF = 1.0

ATF = 0.9

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

160

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

12
40

24.5

GW ~
1000 LB

ATF = 1.0

40
50

50
12
30

24.5

GW ~
1000 LB

30
20

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1065_6
SA

Figure 7A-10. Cruise High Drag - Pressure Altitude Sea Level (Sheet 6 of 6)
7A-31

TM 1-1520-237-10

CRUISE

CRUISE
2,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 2,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
8

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

140

150

30
140

150
130

130

140
120

MAX
RANGE

110

120

110

100
100

80
MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

90
90

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

110

100

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

50
40

40
12

16 18

GW ~
1000 LB

16 18

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-11. Cruise - Pressure Altitude 2,000 Feet (Sheet 1 of 6)
7A-32

AA1061_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
2,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 2,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

160

150

140

140
140

130
MAX
RANGE

130

120

130

MAX
RANGE

120
120

110
110
100

100

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

110

50
50

50
40

12 14

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1061_2
SA

Figure 7A-11. Cruise - Pressure Altitude 2,000 Feet (Sheet 2 of 6)
7A-33

TM 1-1520-237-10

CRUISE

CRUISE
2,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 2,000 FT
0oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

MAX
RANGE

130

MAX
RANGE

130

100

120

110

100

MAX END
AND R / C

100
TRANSMISSION TORQUE LIMIT

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

120

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

120

50
50

40
12

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-11. Cruise - Pressure Altitude 2,000 Feet (Sheet 3 of 6)
7A-34

AA1061_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
2,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 2,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160
10

160

170

30
160

150

150
140

140
MAX
RANGE

130

140

MAX
RANGE

130

130
120

120

110
100

90
MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100
90

80
MAX END
AND R / C
70

100

40
12

GW ~
1000 LB

12 14

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

GW ~
1000 LB
40
30

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1061_4
SA

Figure 7A-11. Cruise - Pressure Altitude 2,000 Feet (Sheet 4 of 6)
7A-35

TM 1-1520-237-10

CRUISE

CRUISE
2,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 2,000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

180
170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

170

150

160

140

150

MAX END
AND R / C
80

110

100

80
MAX END
AND R / C
70

50
12

GW ~
1000 LB

120

GW ~
1000 LB

100

TRUE AIRSPEED ~ KTS

110

130

TRANSMISSION TORQUE LIMIT

100

120

TES ATF = 0.9

110

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

120

140

TORQUE AVAILABLE ~ 30 MINU

130

MAX
RANGE

130

TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE~CONTINUO

140

MAX
RANGE

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-11. Cruise - Pressure Altitude 2,000 Feet (Sheet 5 of 6)
7A-36

AA1061_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
2,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 2,000 FT
60oC

50oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

MAX
RANGE

130

120
TORQUE AVAILABLE ~ 30 MINUTES

130

100

90
MAX END
AND R / C
80

170
160
150
140

130

120

100

110

TRANSMISSION TORQUE LIMIT

110

TRANSMISSION TORQUE LIMIT

120

180

80
MAX END
AND R / C

100

TRUE AIRSPEED ~ KTS

140

ATF = 1.0

140
MAX
RANGE

ATF = 0.9

150

TRUE AIRSPEED ~ KTS

TORQUE AVAILABLE ~ 30 MINUTES

160

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 1.0

170

ATF = 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

60
40

50
12

GW ~
1000 LB

40
30

GW ~
1000 LB

20
10

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 99%

PLACARD TORQUE LIMIT = 96%

DATA BASE: FLIGHT TEST

100

AA1061_6
SA

Figure 7A-11. Cruise - Pressure Altitude 2,000 Feet (Sheet 6 of 6)
7A-37

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T701C (2)

PRESS ALT: 2000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

160

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

160

150
150

150
140

140

140
130

130
130
120
120

MAX
RANGE

120

MAX
RANGE

110

100

100
90

MAX END
AND R / C
70

CONTINUOUS TORQUE LIMIT

80
MAX END
AND R / C
70

TRUE AIRSPEED ~ KTS

100

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110
110

50
40

30
30

24.5

GW ~
1000 LB

24.5

GW ~
1000 LB

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-12. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 1 of 6)
7A-38

AA1062_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T701C (2)

PRESS ALT: 2000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

170

30
160

150

140

130

140

110

100

MAX END
AND R / C

MAX
RANGE

120

110

100

80
MAX END
AND R / C
70

40
12 14 16

24.5

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

CONTINUOUS TORQUE LIMIT

MAX
RANGE

120

130

12 14

24.5

GW ~
1000 LB

20
10
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1062_2
SA

Figure 7A-12. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 2 of 6)
7A-39

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T701C (2)

PRESS ALT: 2000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

180

160

170
10

160

170

150

160

140

150

150
140

140

130
130

130
120

MAX
RANGE

120

MAX
RANGE

110
100

100
100
90

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

80
MAX END
AND R / C

40
12

14 16

24.5

GW ~
1000 LB

50
12

TRUE AIRSPEED ~ KTS

110

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120
110

14 16

24.5
40

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-12. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 3 of 6)
7A-40

AA1062_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T701C (2)

PRESS ALT: 2000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
6

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

180

160
10

170

180
10

30
170

150
160

160
140

150

150
130

120

MAX
RANGE

110

100

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

90
MAX END
AND R / C

130

MAX
RANGE

110

120

100

110
CONTINUOUS TORQUE LIMIT

130

140

80
MAX END
AND R / C

100

50
12

14 16

24.5

GW ~
1000 LB

50
12

TRUE AIRSPEED ~ KTS

140

24.5
40

GW ~
1000 LB

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1062_4
SA

Figure 7A-12. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 4 of 6)
7A-41

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T701C (2)

PRESS ALT: 2000 FT
30oC

40oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR

170

30
~ CONTINUOUS

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

150

170
160

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

140

150

180
170
160
150

130
140

140
MAX
RANGE

130

110

120

90
MAX END
AND R / C

110
90

80
MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

100

TORQUE AVAILABLE ~ 30 MINUTES

110

CONTINUOUS TORQUE LIMIT

TORQUE AVAILABLE
~ CONTINUOUS

120

100

TRUE AIRSPEED ~ KTS

130

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

40
50

GW ~
1000 LB

20
10

~ 30 MINUTE ATF = 0.9

50
24.5

30
12

24.5

GW ~
1000 LB

30
20

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-12. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 5 of 6)
7A-42

AA1062_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
2000 FT
T701C (2)

PRESS ALT: 2000 FT
60oC

50oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

160

150

140

150

MAX
RANGE

130
120
110

100

180
170
160
150

TORQUE AVAILABLE ~ 30 MINUTES

130

140

TRUE AIRSPEED ~ KTS

120
MAX
RANGE
110

100

140
130
120
110

100

90
MAX END
AND R / C

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

160

ATF = 0.9

170

ATF = 1.0

180

ATF = 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

80
60

70
50

60
40

50
12

24.5

GW ~
1000 LB

24.5

GW ~
1000 LB

30
20

10
0

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 99%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1062_6
SA

Figure 7A-12. Cruise High Drag - Pressure Altitude 2,000 Feet (Sheet 6 of 6)
7A-43

TM 1-1520-237-10

CRUISE

CRUISE
4,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 4,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
9

170

160
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

150

160
10

150

140

140
140

130

130
130

MAX
RANGE

120

MAX
RANGE

120

120
110
110
110
100

MAX END
AND R / C
70

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

90
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

100
100

50
50

50
40

14 16

GW ~
1000 LB

14 16

GW ~
1000 LB

30
20

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-13. Cruise - Pressure Altitude 4,000 Feet (Sheet 1 of 6)
7A-44

AA1063_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
4,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 4,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160
10

150

140

130
130

MAX
RANGE

120

110

100

MAX END
AND R / C

100

TRANMISSION TORQUE LIMIT

100

TRANMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

120

80
MAX END
AND R / C
70

40
12

14 16

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

130

MAX
RANGE

GW ~
1000 LB

40
30

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1063_2
SA

Figure 7A-13. Cruise - Pressure Altitude 4,000 Feet (Sheet 2 of 6)
7A-45

TM 1-1520-237-10

CRUISE

CRUISE
4,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 4,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
6

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

160
160

160

150

150
140

140

140
130

MAX
RANGE

130

MAX
RANGE

130

120

110
100

100

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

80
MAX END
AND R / C
70

40
12

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

GW ~
1000 LB

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-13. Cruise - Pressure Altitude 4,000 Feet (Sheet 3 of 6)
7A-46

AA1063_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
4,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 4,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
6

TOTAL FUEL FLOW ~ 100 LB/HR
7

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

180

160
10

10
160
150

170

150

160

140

150

140
TORQUE AVAILABLE ~ CONTINUOUS

130

140

130

MAX
RANGE

120

110
110
100

90
MAX END
AND R / C
80

80
MAX END
AND R / C
70

50
12

GW ~
1000 LB

120

110

100

TRANSMISSION TORQUE LIMIT

100
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

130

TRUE AIRSPEED ~ KTS

180

40
30

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1063_4
SA

Figure 7A-13. Cruise - Pressure Altitude 4,000 Feet (Sheet 4 of 6)
7A-47

TM 1-1520-237-10

CRUISE

CRUISE
4,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 4,000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

~ CONTINUOUS

170

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

160
10

150

180
170
160

140

150

150
MAX
RANGE

130

120

110

100

MAX
RANGE

140

50
12

GW ~
1000 LB

80
MAX END
AND R / C

40
12

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

110
TORQUE AVAILABLE ~ 30 MINUTES ATF= 0.9

ATF= 0.9

100

TRANSMISSION TORQUE LIMIT

120

100

TRUE AIRSPEED ~ KTS

130

TORQUE AVAILABLE ~ 30 MINUTES ATF= 0.9

TRUE AIRSPEED ~ KTS

140

130

20
10

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT = 94%

DATA BASE: FLIGHT TEST

Figure 7A-13. Cruise - Pressure Altitude 4,000 Feet (Sheet 5 of 6)
7A-48

100

AA1063_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
4,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 4,000 FT
60oC

50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

160
10

150

160

ATF= 1.0

ATF= 0.9

~CONTINUOUS

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 1.0

~CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF= 0.9

170

180
170
160

140

150
140

150

130

MAX
RANGE

140

120

110

100

80
MAX END
AND R / C

TORQUE AVAILABLE ~ 30 MINUTES

100

60
MAX END
AND R / C
50

110

100

TRUE AIRSPEED ~ KTS

130
110

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

130

60
40

GW ~
1000 LB

50
12

40
20

10
10

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 91%

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMNIT = 88%

DATA BASE: FLIGHT TEST

AA1063_6
SA

Figure 7A-13. Cruise - Pressure Altitude 4,000 Feet (Sheet 6 of 6)
7A-49

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T701C (2)

PRESS ALT: 4000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

170

160
160

160

150

150
140

130

120

130

120
MAX
RANGE

110

120

MAX
RANGE
110

100

110

100

MAX END
AND R / C

100
CONTINUOUS TORQUE LIMIT

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

80
MAX END
AND R / C
70

TRUE AIRSPEED ~ KTS

140

50
50

40
40

GW ~
1000 LB

14 16 18

24.5

14 16

24.5

GW ~
1000 LB

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-14. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 1 of 6)
7A-50

AA1064_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T701C (2)

PRESS ALT: 4000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
7

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

170

160

180

170

150

160

150

140
140

140
130

130

130
120
MAX
RANGE

120

MAX
RANGE

120

110
100

100

100
90

90
MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

80
MAX END
AND R / C

40
12

24.5

TRUE AIRSPEED ~ KTS

110

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

24.5
40

GW ~
1000 LB

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1064_2
SA

Figure 7A-14. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 2 of 6)
7A-51

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T701C (2)

PRESS ALT: 4000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

180

170

150
160

160
140

150

150
130

140
130

120

130

MAX
RANGE

110

CONTINUOUS TORQUE LIMIT

100

90
MAX END
AND R / C

110

120

100

110

100

80
MAX END
AND R / C

50
12

24.5

24.5
40

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

MAX
RANGE

CONTINUOUS TORQUE LIMIT

120

TRUE AIRSPEED ~ KTS

140

10
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-14. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 3 of 6)
7A-52

AA1064_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T701C (2)

PRESS ALT: 4000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
20

160

TORQUE AVAILABLE
~ CONTINUOUS

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
30

150

170
160

140

180
170
160

150

130
140

140
120
130

MAX
RANGE

110

120

120
100

110

110
90

100

90
MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

100
80
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

MAX
RANGE

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

40
24.5
12

24.5

GW ~
1000 LB

30
20

10
0

50
~ 30 MINUTES ATF = 0.9

40
30
20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1064_4
SA

Figure 7A-14. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 4 of 6)
7A-53

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T701C (2)

PRESS ALT: 4000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

30
180

150

140

130

150
140

170
160
150

120
MAX
RANGE

130

140
MAX
RANGE

110

130

120

100
110

110
90

80
90
70

MAX END
AND R / C

24.5

MAX END
AND R / C

50
24.5
40

30
40

GW ~
1000 LB

TORQUE AVAILABLE ~ 30 MINUTES

100

TORQUE AVAILABLE
~ 30 MINUTES

TRUE AIRSPEED ~ KTS

ATF = 1.0

160

ATF = 0.9

170

ATF = 0.9

180

100

GW ~
1000 LB

30
20

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

30
20

0
20

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 97%

100

0
20

100

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 94%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-14. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 5 of 6)
7A-54

AA1064_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
4000 FT
T701C (2)

PRESS ALT: 4000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

13
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

ATF = 1.0

150

TRUE AIRSPEED ~ KTS

140

MAX
RANGE

130
120

140

130

120

110

100

110

100

90
70
80

MAX END
AND R / C

INDICATED AIRSPEED ~ KTS

160

ATF = 0.9

~ CONTINUOUS

170

TORQUE AVAILABLE ~ 30 MINUTES

150

180

70
50
60
40

24.5
50
12

40
GW ~
1000 LB

10
0

0
20

100

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 91%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1064_6
SA

Figure 7A-14. Cruise High Drag - Pressure Altitude 4,000 Feet (Sheet 6 of 6)
7A-55

TM 1-1520-237-10

CRUISE

CRUISE
6,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 6,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160

150
150

150

140

130

140

130
MAX
RANGE

120

MAX
RANGE

120

100

MAX END
AND R / C

80
MAX END
AND R / C
70

40
40

12 14

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110

TRANSMISSION TORQUE LIMIT

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

GW ~
1000 LB

20
20

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-15. Cruise - Pressure Altitude 6,000 Feet (Sheet 1 of 6)
7A-56

AA1009_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
6,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 6,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

140
130

130

MAX
RANGE

130

120

110

90
MAX END
AND R / C

GW ~
1000 LB

100

80
MAX END
AND R / C
70

100

50
12

TRANSMISSION TORQUE LIMIT

100

TRUE AIRSPEED ~ KTS

110
TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

110

GW ~
1000 LB

40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1009_2
SA

Figure 7A-15. Cruise - Pressure Altitude 6,000 Feet (Sheet 2 of 6)
7A-57

TM 1-1520-237-10

CRUISE

CRUISE
6,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 6,000 FT
0o C

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

180
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

160
150

180

170

150

160

140

150

130

140

140
MAX
RANGE

120

130

110

120

100

110

100

90
MAX END
AND R / C
80

TRANSMISSION TORQUE LIMIT

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

80
MAX END
AND R / C
70

50
12

GW ~
1000 LB

50
12

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

GW ~
1000 LB
40
30

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-15. Cruise - Pressure Altitude 6,000 Feet (Sheet 3 of 6)
7A-58

AA1009_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
6,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 6,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

180

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

170
150
160
140

170
160

150

150
130

140
MAX
RANGE

130

90
MAX END
AND R / C

50
12

GW ~
1000 LB

40
30
20
10

TORQUE AVAILABLE ~ 30 MINUTES ATF
= 0.9

110
90

80
MAX END
AND R / C

40
12

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

100

100
TORQUE AVAILABLE ~ 30 MINUTES ATF =
0.9

~ CONTINUOUS

110

120

100

50
40

30
20

10
10

0
20

TRUE AIRSPEED ~ KTS

110

120

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

140

MAX
RANGE

120

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT =93%

DATA BASE: FLIGHT TEST

100

AA1009_4
SA

Figure 7A-15. Cruise - Pressure Altitude 6,000 Feet (Sheet 4 of 6)
7A-59

TM 1-1520-237-10

CRUISE

CRUISE
6,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 6,000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

160

140

150
130
MAX
RANGE

140

MAX
RANGE

120

130
110
120

180
170
160
150
140

130

120

100
110

110

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

90
TRANSMISSION TORQUE LIMIT

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

MAX END
AND R / C

100

TRUE AIRSPEED ~ KTS

6
170

ATF = 1.0

TES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
180

TORQUE AVAILABLE ~ 30 MINU

ATF = 0.9

TOTAL FUEL FLOW ~ 100 LB/HR

~ CONTINUOUS

ATF = 1.0

ATF = 0.9

~ CONTINUOUS

IAS ~ KTS

40
50
12

GW ~
1000 LB

40
20

20
10

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 90%

PLACARD TORQUE LIMIT = 87%

DATA BASE: FLIGHT TEST

Figure 7A-15. Cruise - Pressure Altitude 6,000 Feet (Sheet 5 of 6)
7A-60

100

AA1009_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
6,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 6,000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

~ CONTINUOUS

180
170
160

MAX
RANGE

130

TRUE AIRSPEED ~ KTS

150

140

150
140

160

120

110

130

120

110

100

INDICATED AIRSPEED ~ KTS

ATF = 1.0

TORQUE AVAILABLE ~ 30 MINUTES

ATF = 0.9

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
MAX END
AND R / C

50
60
40
50

GW ~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 83%

DATA BASE: FLIGHT TEST

AA1009_6
SA

Figure 7A-15. Cruise - Pressure Altitude 6,000 Feet (Sheet 6 of 6)
7A-61

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T701C (2)

PRESS ALT: 6000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
20

30
~ CONTINUOUS

160

160
150
140

180

170
~ CONTINUOUS

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

150

140

160
150
140

130
130

130
120
120

110
110

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

100

80
MAX END
AND R / C

100

60
60

14 16

GW ~
1000 LB

24.5

14 16

24.5

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

110

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

MAX
RANGE

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-16. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 1 of 6)
7A-62

AA1010_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T701C (2)

PRESS ALT: 6000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

30
~ CONTINUOUS

170
160
150

130

120
MAX
RANGE

110

MAX END
AND R / C

170
160
150
140
130

MAX
RANGE

110

100

180

140

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

150

140

120

100

80
MAX END
AND R / C

110

100

TRUE AIRSPEED ~ KTS

CONTINUOUS TORQUE LIMIT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
~ CONTINUOUS

180

50
12

14 16

24.5

GW ~
1000 LB

24.5

GW ~
1000 LB

20
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1010_2
SA

Figure 7A-16. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 2 of 6)
7A-63

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T701C (2)

PRESS ALT: 6000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

180

160

30
180

150

170

160

140

160

130

150

150
140

140
120
130

MAX
RANGE

110

120

120
100

110

100

90
MAX END
AND R / C

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

110
CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

MAX
RANGE

100

TRUE AIRSPEED ~ KTS

130

50
60

60
40

50
12

24.5

24.5
30

GW ~
1000 LB

40
GW ~
1000 LB

30
20

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-16. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 3 of 6)
7A-64

AA1010_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T701C (2)

PRESS ALT: 6000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
160

140

150
140

120
MAX
RANGE

MAX
RANGE

110

120
100
110
90

ATF = 0.9

100

MAX END
AND R / C

24.5

50
12

40
GW ~
1000 LB

30
20

180
170
160
150
140
130
120
110

100
80
90
70
80
MAX END
AND R / C

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

150

130

TORQUE AVAILABLE ~ 30 MINU

180

TRUE AIRSPEED ~ KTS

160

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TES

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

ATF = 0.9

170

70
50
60
40

24.5

GW ~
1000 LB

40
30
20

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT = 93%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1010_4
SA

Figure 7A-16. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 4 of 6)
7A-65

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T701C (2)

PRESS ALT: 6000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

120
110
100
90

ATF = 1.0

150

120
MAX
RANGE

110

160

100

140
130
120
110
100

80
MAX END
AND R / C

MAX END
AND R / C

24.5

50
12

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

MAX
RANGE

170

TORQUE AVAILABLE ~ 30 MINUTES

140

ATF = 0.9

140

130

150

130

180

~ CONTINUOUS

160

ATF = 1.0

~ CONTINUOUS

170

ATF = 0.9

150

180

TRUE AIRSPEED ~ KTS

40
GW ~
1000 LB

30
20

0
20

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 90%

100

0
20

100

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 87%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-16. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 5 of 6)
7A-66

AA1010_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
6000 FT
T701C (2)

PRESS ALT: 6000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

13
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

130

TRUE AIRSPEED ~ KTS

140
MAX
RANGE

130
120
110
100

120

110

100

90
MAX END
AND R / C

INDICATED AIRSPEED ~ KTS

150

ATF = 1.0

160

140

TORQUE AVAILABLE ~ 30 MINUTES

170

ATF = 0.9

~ CONTINUOUS

180

50
24.5

GW ~
1000 LB

10
0

0
20

100

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 83%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASIS:FLIGHT TEST

AA2101_6
SA

Figure 7A-16. Cruise High Drag - Pressure Altitude 6,000 Feet (Sheet 6 of 6)
7A-67

TM 1-1520-237-10

CRUISE

CRUISE
8,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 8,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR

170
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

160

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170

30
160

150

150
140

140

110

100
100

MAX END
AND R / C

TORQUE AVAILABLE

GW ~
1000 LB

120

110

100

90
80
MAX END
AND R / C

TORQUE AVAILABLE

130

TRUE AIRSPEED ~ KTS

MAX
RANGE

110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

MAX
RANGE

120

TRANSMISSION TORQUE LIMIT

130

40
12

GW ~
1000 LB
40
30

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-17. Cruise - Pressure Altitude 8,000 Feet (Sheet 1 of 6)
7A-68

AA1011_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 8000 FT
−30OC

−20OC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

160

170

150

160

140

150

130

140

150

140

130
MAX
RANGE

MAX
RANGE

120

130

120

100

110

100

90
MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

100

80
MAX END
AND R / C

50
12

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

110
110

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

120

GW ~
1000 LB
40
30

20
10
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

AA1011_2
SA

Figure 7A-17. Cruise - Pressure Altitude 8,000 Feet (Sheet 2 of 6)
7A-69

TM 1-1520-237-10

CRUISE

CRUISE
8,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 8,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

160

180

170
170
150
160

160
140

150
150
130

140

TORQUE AVAILABLE
~ CONTINUOUS

120

90
MAX END
AND R / C
80

80
MAX END
AND R / C

40
50
12

GW ~
1000 LB

110

100
TRANSMISSION TORQUE LIMIT

100

50
ATF = 1.0

100

110

~ CONTINUOUS

110

130

TORQUE AVAILABLE
~30 MINUTES ATF = 0.9

120

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

MAX
RANGE

120

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

20
10

10
0

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT = 93%

DATA BASE: FLIGHT TEST

Figure 7A-17. Cruise - Pressure Altitude 8,000 Feet (Sheet 3 of 6)
7A-70

100

AA1011_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
8,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 8,000 FT
20oC

10oC

170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

160

150
130
140
120

MAX
RANGE

130
110
120
100
110

160

140

150

130

120

110

90
100

100
80

50
12
40

GW ~
1000 LB

MAX END
AND R / C

40
12

50
40

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

180
170

140

MAX
RANGE

170

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
180

ATF = 1.0

30 MINUTES

TORQUE AVAILABLE ~

TOTAL FUEL FLOW ~ 100 LB/HR

ATF = 0.9

ATF = 0.9
ATF = 1.0

~ CONTINUOUS

IAS ~ KTS

~ CONTINUOUS

TOTAL FUEL FLOW ~ 100 LB/HR

30
20

20
10

10
0

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 90%

PLACARD TORQUE LIMIT = 86%

DATA BASE: FLIGHT TEST

100

AA1011_4
SA

Figure 7A-17. Cruise - Pressure Altitude 8,000 Feet (Sheet 4 of 6)
7A-71

TM 1-1520-237-10

CRUISE

CRUISE
8,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 8,000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

130
120

100

90
MAX END
AND R / C
80
70

MAX
RANGE

110

100

TRANSMISSION TORQUE LIMIT

110

170

130

120

180

160
150
140
130
120

110

100

TRUE AIRSPEED ~ KTS

MAX
RANGE

140

ATF = 1.0

TORQUE AVAILABLE ~ 30 MINUTES

150

TRUE AIRSPEED ~ KTS

150

160

140

ATF = 0.9

170

160

~ CONTINUOUS

180

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TORQUE AVAILABLE ~ 30 MINUTES

~ CONTINUOUS

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MAX END
AND R / C
70

60
40

GW ~
1000 LB

30
20

10
0

0
20

100

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 83%

PLACARD TORQUE LIMIT = 81%

DATA BASE: FLIGHT TEST

Figure 7A-17. Cruise - Pressure Altitude 8,000 Feet (Sheet 5 of 6)
7A-72

AA1011_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
8,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 8,000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

12
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

170
160
150

MAX
RANGE

140

TRUE AIRSPEED ~ KTS

150
TORQUE AVAILABLE ~ 30 MINUTES

180

130
120

140

130

120

110

100

90
110
80

100

90
80

MAX END
AND R / C

40
12

GW ~
1000 LB

INDICATED AIRSPEED ~ KTS

ATF = 0.9

ATF = 1.0

~ CONTINUOUS

30
20

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 77%

DATA BASE: FLIGHT TEST

AA1011_6
SA

Figure 7A-17. Cruise - Pressure Altitude 8,000 Feet (Sheet 6 of 6)
7A-73

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

PRESS ALT: 8000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

180

160

180

170
170

150

160

160
140

150

150
130

130

120

130

110

MAX
RANGE
CONTINUOUS TORQUE LIMIT

110

100

MAX END
AND R / C

120

MAX
RANGE

100

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

120

110

100

24.5

GW ~
1000 LB

30
20
10

50
12 14

TORQUE AVAILABLE

12 14

24.5

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

140

GW ~
1000 LB
20

40
30
20
10
0

0
20

TORQUE PER ENGINE ~ %

100

TRUE AIRSPEED ~ KTS

140

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-18. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 1 of 6)
7A-74

AA1012_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

PRESS ALT: 8000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

30
180

150

170

170
160

140

160

150
150

130
140

140
120

MAX
RANGE

120

MAX
RANGE

120

100

110

110
90

100

MAX END
AND R / C

CONTINUOUS TORQUE LIMIT

100

TRUE AIRSPEED ~ KTS

130
110

CONTINUOUS TORQUE LIMIT

TRUE AIRSPEED ~ KTS

130

60
40

24.5

40
12
30

GW ~
1000 LB

24.5

GW ~
1000 LB

10
10

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1012_2
SA

Figure 7A-18. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 2 of 6)
7A-75

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

PRESS ALT: 8000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
6

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160
30

180

150

170

OUS

140

ABLE ~ CONTINU

150
140

MAX
RANGE

120
110

160

130
150
120

140

110

TORQUE AVAIL

TRUE AIRSPEED ~ KTS

130

170

130

MAX
RANGE

120

100

110

ATF= 1.0

160

180

100
80
90

100
90

70
80

80
60

MAX END
AND R / C

50
60
24.5

GW ~
1000 LB

40
30
20
10

~ 30 MINUTE ATF= 0.9

24.5

40
12

GW ~
1000 LB

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT = 93%

TORQUE AVAILABLE ~ 30 MINUTES

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

ATF= 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

50
40
30
20
10

0
100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-18. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 3 of 6)
7A-76

AA1012_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

PRESS ALT: 8000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

150

140

160

180
170
160

130
~ 30 MINUTES

150
120

110

120
110
100
90

MAX END
AND R / C

MAX
RANGE

TORQUE AVAILABLE

MAX
RANGE

130

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

140

100

150
140
130
120
110

100

TRUE AIRSPEED ~ KTS

170

ATF = 1.0

180

160

ATF = 0.9

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
ATF = 0.9
ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

24.5
70

50
MAX END
AND R / C

40
24.5
50
23
40

GW ~
1000 LB

10
0

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 90%

PLACARD TORQUE LIMIT = 86%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1012_4
SA

Figure 7A-18. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 4 of 6)
7A-77

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

PRESS ALT: 8000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
7

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

130

ATF= 1.0

140

160

180
170

MAX
RANGE

130
120
110
100

120

110
MAX
RANGE

100

150
140
130
120
110
100
90

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

140

TORQUE AVAILABLE ~ 30 MINUTES

160

150

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

170

ATF= 1.0

180

150

ATF= 0.9

30
ATF= 0.9

~ CONTINUOUS

80
MAX END
AND R / C

70
23

40
12

60
12

GW ~
1000 LB

20
30

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 83%

PLACARD TORQUE LIMIT = 81%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-18. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 5 of 6)
7A-78

AA1012_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
8000 FT
T701C (2)

PRESS ALT: 8000 FT
50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

12
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

140

130

TRUE AIRSPEED ~ KTS

140
MAX
RANGE

130

120

120
110
100

110

100

70
90

INDICATED AIRSPEED ~ KTS

150

ATF = 1.0

160

TORQUE AVAILABLE ~ 30 MINUTES

170

ATF = 0.9

~ CONTINUOUS

180

60
80
MAX END
AND R / C

60
22

50
12

20
20

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 77%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASIS:FLIGHT TEST

AA2102_6
SA

Figure 7A-18. Cruise High Drag - Pressure Altitude 8,000 Feet (Sheet 6 of 6)
7A-79

TM 1-1520-237-10

CRUISE

CRUISE
10,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 10,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

170

160

170

160
150

160

140

150

130

140

120

130

150

140

130

MAX END
AND R / C

100

50
12
40

50
12

22
40

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

NUOUS

110

TRANSMISSION TORQUE LIMIT

30 MINUTES & CONTI

MAX END
AND R / C

TORQUE AVAILABLE~

100

TRANSMISSION TORQUE LIMIT

100

120

110

NUOUS
30 MINUTES & CONTI

110

TORQUE AVAILABLE~

TRUE AIRSPEED ~ KTS

MAX
RANGE

120

20
10

10
0

0
20

TORQUE PER ENGINE ~ %

100

TORQUE PER ENGINE ~ %

DATA BASE: FLIGHT TEST

Figure 7A-19. Cruise - Pressure Altitude 10,000 Feet (Sheet 1 of 5)
7A-80

AA1013_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
10,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 10,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

160

180

170
170
150
160
160
140

150

150
130

140

140
120
MAX
RANGE

130

110

120

TORQUE AVAILABLE~

50
12
40

MAX END
AND R / C

40
12

30
GW ~
1000 LB

GW ~
1000 LB

TRANSMISSION TORQUE LIMIT

MAX END
AND R / C

110

100

ATF = 0.9

~ CONTINUO
US

100

TRANSMISSION TORQUE LIMIT

110

TORQUE AVAILABLE ~ 30 MINUTES

NUOUS

120

TRUE AIRSPEED ~ KTS

MAX
RANGE

30 MINUTES & CONTI

TRUE AIRSPEED ~ KTS

130

20
10

10
0

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT = 93%

DATA BASE: FLIGHT TEST

100

AA1013_2
SA

Figure 7A-19. Cruise - Pressure Altitude 10,000 Feet (Sheet 2 of 5)
7A-81

TM 1-1520-237-10

CRUISE

CRUISE
10,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 10,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
6

170

150

160

170

TORQUE AVAILABLE ~ 30 MINU

140

150

130

120
MAX
RANGE
110

120
100

160
150
140

130

120

90
MAX END
AND R / C

50
12

70
MAX END
AND R / C
60

100

60
40
12

GW ~
1000 LB

110

TRANSMISSION TORQUE LIMIT

TES

100

TRANSMISSION TORQUE LIMIT

110

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

ATF = 1.0

140
MAX
RANGE

180

TRUE AIRSPEED ~ KTS

ATF = 0.9

ATF = 1.0

TES

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 0.9

~ CONTINUOUS

170

10
10

10
0

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 89%

PLACARD TORQUE LIMIT = 85%

DATA BASE: FLIGHT TEST

Figure 7A-19. Cruise - Pressure Altitude 10,000 Feet (Sheet 3 of 5)
7A-82

100

AA1013_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
10,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 10,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

130

120

110

100

90
MAX END
AND R / C

80
70

TES

ATF = 0.9

ATF = 1.0

170

120
MAX
RANGE
110

100

70
MAX END
AND R / C

160
150
140
130
120

110

100
90
80
70

40
50

GW ~
1000 LB

30
20

TRUE AIRSPEED ~ KTS

MAX
RANGE

130

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

180

TRANSMISSION TORQUE LIMIT

150

TORQUE AVAILABLE ~ 30 MINU

160

150

TORQUE AVAILABLE ~ 30 MINU

170

~ CONTINUOUS

ATF = 1.0

180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

TES

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

ATF = 0.9

170

30
20

10
0

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 83%

PLACARD TORQUE LIMIT = 80%

DATA BASE: FLIGHT TEST

100

AA1013_4
SA

Figure 7A-19. Cruise - Pressure Altitude 10,000 Feet (Sheet 4 of 5)
7A-83

TM 1-1520-237-10

CRUISE

CRUISE
10,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 10,000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MAX
RANGE

130
120

110

TES

170

120
MAX
RANGE

110

160

100

150
140
130
120

110

80
100

100
70

TRUE AIRSPEED ~ KTS

130

180

TORQUE AVAILABLE ~ 30 MINU

150
140

140

TORQUE AVAILABLE ~ 30 MINU

160

ATF = 1.0

170

150

TES

180

ATF = 1.0

ATF = 0.9

~ CONTINUOUS

160

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
60

MAX END
AND R / C

GW ~
1000 LB

GW ~ 12
1000 LB

40
20

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 77%

PLACARD TORQUE LIMIT = 74%

DATA BASE: FLIGHT TEST

Figure 7A-19. Cruise - Pressure Altitude 10,000 Feet (Sheet 5 of 5)
7A-84

AA1013_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T701C (2)

PRESS ALT: 10000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
20

160

~ CONTINUOUS

170
160
150

30
180

~ CONTINUOUS

180

150

140

170
160
150

130
140

140
120

MAX
RANGE
110

100

80
MAX END
AND R / C
70

110
MAX
RANGE

100

120
110

100

TRUE AIRSPEED ~ KTS

120

130
TORQUE AVAILABLE ~ 30 MINUTES

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

130

MAX END
AND R / C

24.5
24.5

23
12

30
40

GW ~
1000 LB

30
20

10
0

0
20

TORQUE PER ENGINE ~ %

100

0
20

100

TORQUE PER ENGINE ~ %

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1014_1
SA

Figure 7A-20. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 1 of 5)
7A-85

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T701C (2)

PRESS ALT: 10000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

160

180

150

AND CONTINUOU

170

150
120

MAX
RANGE

110

100

140

110

130
MAX
RANGE

100

TORQUE AVAILAB

TRUE AIRSPEED ~ KTS

130
120

160
130

LE ~ 30 MINUTES

140

80
60
70

170

30 MINUTES

150

140

TORQUE AVAILABLE ~

160

180

MAX END
AND R / C

120
110
100

TRUE AIRSPEED ~ KTS

~ CONTINUO

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

70
50

MAX END
AND R / C

24.5

60
24.5

50
23

40
12

GW ~
1000 LB

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 97%

PLACARD TORQUE LIMIT = 93%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-20. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 2 of 5)
7A-86

AA1014_2
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T701C (2)

PRESS ALT: 10000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

160

150

140

130

170
160

150

30 MINUTES

120

110
MAX
RANGE

120
110
100

MAX
RANGE

100

TORQUE AVAILABLE ~

TORQUE AVAILABLE
~ 30 MINUTES

TRUE AIRSPEED ~ KTS

140
130

180

140
130
120
110
100
90

TRUE AIRSPEED ~ KTS

170

ATF = 0.9
ATF = 1.0

180

160

~ CONTINUOUS

ATF = 0.9
ATF = 1.0

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

24.5
70

MAX END
AND R / C

40
23
50

23
12

GW ~
1000 LB

40
30

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 89%

PLACARD TORQUE LIMIT = 85%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1014_3
SA

Figure 7A-20. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 3 of 5)
7A-87

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T701C (2)

PRESS ALT: 10000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

130

TES

120

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

MAX
RANGE

130
120
110
100

180
170
160

110
MAX
RANGE

100

ATF = 1.0

150

140

150
140

150

140
130
120
110
100

TRUE AIRSPEED ~ KTS

160

TORQUE AVAILABLE ~ 30 MINU

170

ATF = 0.9
ATF = 1.0

180

TES

160

~ CONTINUOUS

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

90
60

MAX END
AND R / C

23
50

23
12

GW ~
1000 LB

30
20

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 83%

PLACARD TORQUE LIMIT = 80%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-20. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 4 of 5)
7A-88

AA1014_4
SA

TM 1-1520-237-10

CRUISE

CRUISE
10000 FT
T701C (2)

PRESS ALT: 10000 FT
40oC

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

TRUE AIRSPEED ~ KTS

140
MAX
RANGE

120
110
100

ATF = 1.0

ATF = 0.9

170

120

110
MAX
RANGE

100

180

160
150
140
130
120
110
100

70
90

TRUE AIRSPEED ~ KTS

130

150

130

140

TORQUE AVAILABLE ~ 30 MINUTES

160

TORQUE AVAILABLE ~ 30 MINUTES

170

~ CONTINUOUS

180

150

ATF = 1.0

ATF = 0.9

~ CONTINUOUS

90
60

80
MAX END
AND R / C

MAX END
AND R / C

60
12
50

12
30

GW ~
1000 LB

20
50

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 77%

PLACARD TORQUE LIMIT = 74%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1014_5
SA

Figure 7A-20. Cruise High Drag - Pressure Altitude 10,000 Feet (Sheet 5 of 5)
7A-89

TM 1-1520-237-10

CRUISE

CRUISE
12,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 12,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
12

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

140

TORQUE AVAILAB
LE ~

TRUE AIRSPEED ~ KTS

130
MAX
RANGE

120

110

100

90
MAX END
AND R / C

30 MINUTES & CO
NTINUOUS

150

140

130

120

MAX
RANGE

110

100

70
MAX END
AND R / C
60

170
160
150
140

130

120

110
TRANSMISSION TORQUE LIMIT

160

TORQUE AVAILAB
LE ~

TRANSMISSION TORQUE LIMIT

30 MINUTES & CO
NTINUOUS

160
170

100

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

60
40

50
12

GW ~
1000 LB

22
40

GW ~
1000 LB
20

20
10

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 98%

PLACARD TORQUE LIMIT = 93%

DATA BASE: FLIGHT TEST

Figure 7A-21. Cruise - Pressure Altitude 12,000 Feet (Sheet 1 of 5)
7A-90

100

AA1015_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
12,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 12,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

140

TORQUE AVAILAB

MAX
RANGE

140

120

110

100

90
MAX END
AND R / C

130

120
MAX
RANGE

110

180
170
160
150
140

130

120

100

110

70
MAX END
AND R / C

100

TRUE AIRSPEED ~ KTS

150

LE ~ 30 MINUTES

150

TORQUE AVAILAB

160

TRANSMISSION TORQUE LIMIT

170

130

& CONTINUOUS

TRANSMISSION TORQUE LIMIT

LE ~ 30 MINUTES

160

180

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

60
40

50
12

30
40
12
30

GW ~
1000 LB

10
10

10
0

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 89%

PLACARD TORQUE LIMIT = 86%

DATA BASE: FLIGHT TEST

100

AA1015_2
SA

Figure 7A-21. Cruise - Pressure Altitude 12,000 Feet (Sheet 2 of 5)
7A-91

TM 1-1520-237-10

CRUISE

CRUISE
12,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 12,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
6

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170

120

100

110

100

90
MAX END
AND R / C

80
70

MAX
RANGE

110

MAX END
AND R / C

40
12

150
140
130
120

100
90
80
70
60

GW ~
1000 LB

1.0

160

110

GW ~
1000 LB

170

TRANSMISSION TORQUE LIMIT

MAX
RANGE

180

TRUE AIRSPEED ~ KTS

130

TRANSMISSION TORQUE LIMIT

TRUE AIRSPEED ~ KTS

140

130

140

TORQUE AVAILABLE

150

~ 30 MINUTES ATF=

1.0

160

TORQUE AVAILABLE

170

~ CONTINUOUS

180

160

ATF= 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ 30 MINUTES ATF=

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

40
30
20

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 82%

PLACARD TORQUE LIMIT = 79%

DATA BASE: FLIGHT TEST

Figure 7A-21. Cruise - Pressure Altitude 12,000 Feet (Sheet 3 of 5)
7A-92

100

AA1015_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
12,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 12,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

MAX
RANGE

120

110

ATF= 1.0

ATF= 0.9

170

TES

130

120

110

160

MAX
RANGE

100

150
140
130
120

110

80
100

100
70

TRUE AIRSPEED ~ KTS

140

180

TORQUE AVAILABLE ~ 30 MINU

MINUTES
TORQUE AVAILABLE ~ 30

150

TRUE AIRSPEED ~ KTS

150

140

160

130

~ CONTINUOUS

170

ATF= 1.0

180

10
ATF= 0.9

~ CONTINUOUS

90
60

MAX END
AND R / C

70
60

GW ~
1000 LB

MAX END
AND R / C

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 76%

PLACARD TORQUE LIMIT = 73%

DATA BASE: FLIGHT TEST

AA1015_4
SA

Figure 7A-21. Cruise - Pressure Altitude 12,000 Feet (Sheet 4 of 5)
7A-93

TM 1-1520-237-10

CRUISE

CRUISE
12,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 12,000 FT

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

11
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

180

150
ATF= 1.0

ATF= 0.9

~ CONTINUOUS

160
10

140

170
130
TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

150
140

MAX
RANGE

130
120
110

120

110

100

100
70

INDICATED AIRSPEED ~ KTS

160

90
60
MAX END
AND R / C

60
22
12

GW ~
1000 LB

40
20
30
20

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 70%

DATA BASE: FLIGHT TEST

Figure 7A-21. Cruise - Pressure Altitude 12,000 Feet (Sheet 5 of 5)
7A-94

AA1015_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T701C (2)

PRESS ALT: 12000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

13 14

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

180

110

100

MAX
RANGE

100

TORQUE AVAIL

MAX
RANGE

110

180
170
160
150
140
130
120
110
100

90
70
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

120

ABLE ~ 30 MINU

130
120

130

ABLE ~ 30 MINU

140

TES AND CONT

140

TORQUE AVAIL

150

INUOUS

160

TRUE AIRSPEED ~ KTS

150

TES AND CONT

170

80
60

MAX END
AND R / C

24.5

60
40

50
12

GW ~
1000 LB

40
30
20

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 98%

PLACARD TORQUE LIMIT = 93%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1016_1
SA

Figure 7A-22. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 1 of 5)
7A-95

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T701C (2)

PRESS ALT: 12000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

ABLE ~ 30 MINU

130
MAX
RANGE

120
110
100
90

110
MAX
RANGE

100

160

120

TORQUE AVAIL

TRUE AIRSPEED ~ KTS

140

170

~ 30 MINUTES

150

130

180

MAX END
AND R / C

150
140
130
120
110
100

80
70

50
MAX END
AND R / C

40
12

30
GW ~
1000 LB

TRUE AIRSPEED ~ KTS

TES AND CONT

160

140

TORQUE AVAILABLE

170

~ CONTINUOUS

INUOUS

180

GW ~
1000 LB

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMT = 89%

PLACARD TORQUE LIMIT = 86%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-22. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 2 of 5)
7A-96

AA1016_2
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T701C (2)

PRESS ALT: 12000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170

130
MAX
RANGE

120

110

170

110

TORQUE AVAIL
ABLE

TRUE AIRSPEED ~ KTS

140

130

100

180

160
150
140

MAX
RANGE

130

80
100

120
110
100

TRUE AIRSPEED ~ KTS

150

140

TORQUE AVAILAB

160

ATF = 0.9

170

150

~ 30 MINUTES

~ CONTINUOUS

180

ATF = 0.9
ATF = 1.0

160

~ CONTINUOUS

ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

LE ~ 30 MINUTE

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

50
MAX END
AND R / C

80
70
MAX END
AND R / C

23
50

GW ~
1000 LB

50
22
12

GW ~
1000 LB

30
20

10
0

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 82%

PLACARD TORQUE LIMIT = 79%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1016_3
SA

Figure 7A-22. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 3 of 5)
7A-97

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T701C (2)

PRESS ALT: 12000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

120
110

ATF = 0.9

110

160
150

100

140

MAX
RANGE
90

130
120
110

100

100
70

TRUE AIRSPEED ~ KTS

MAX
RANGE

180
170

120

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

140

ATF = 1.0

ATF = 0.9
ATF = 1.0

130

150

130

140

~ 30 MINUTES

160

TORQUE AVAILABLE

170

150

~ 30 MINUTES

180

~ CONTINUOUS

90
60

80
MAX END
AND R / C

50
70
MAX END
AND R / C

40
60
30

50
12

40
30

50
12

GW ~
1000 LB

30
20

10
0

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 76%

PLACARD TORQUE LIMIT = 73%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-22. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 4 of 5)
7A-98

AA1016_4
SA

TM 1-1520-237-10

CRUISE

CRUISE
12000 FT
T701C (2)

PRESS ALT: 12000 FT
30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

12
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
20

160

ATF = 1.0

180

150
ATF = 0.9

~ CONTINUOUS

140

130

TES

170

140
MAX
RANGE

120
110

110

100

INDICATED AIRSPEED ~ KTS

TRUE AIRSPEED ~ KTS

150

130

120

TORQUE AVAILABLE ~ 30 MINU

160

60
80
MAX END
AND R / C

70
40
60
12
50

20
30

GW ~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 70%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1016_5
SA

Figure 7A-22. Cruise High Drag - Pressure Altitude 12,000 Feet (Sheet 5 of 5)
7A-99

TM 1-1520-237-10

CRUISE

CRUISE
14,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 14,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
180

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

30
180

150

170

110
MAX
RANGE
100

TORQUE AVAILAB

110

100

140

130

120

110

100

70
80

TRUE AIRSPEED ~ KTS

MAX
RANGE

150

LE ~ 30 MINUTES

130

120

LE ~ 30 MINUTES

140

130

160

TORQUE AVAILAB

150

TRUE AIRSPEED ~ KTS

140

& CONTINUOUS

160

& CONTINUOUS

170

80
60
MAX END
AND R / C

MAX END
AND R / C

50
60

60
40

50
12
40

GW ~
1000 LB

10
10

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 90%

PLACARD TORQUE LIMIT = 86%

DATA BASE: FLIGHT TEST

Figure 7A-23. Cruise - Pressure Altitude 14,000 Feet (Sheet 1 of 5)
7A-100

AA1017_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
14,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 14,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170

160

180

150

LE ~ 30 MINUTES

130

MAX
RANGE

120

110

100

130

120

110
MAX
RANGE
100

TORQUE AVAILAB

TRUE AIRSPEED ~ KTS

140

170
30 MINUTES

160

140

180

160
150

TORQUE AVAILABLE ~

& CONTINUOUS

170

140
130
120

110

100

MAX END
AND R / C

22
50

GW ~
1000 LB

30
20

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

30
20

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 82%

PLACARD TORQUE LIMIT = 79%

DATA BASE: FLIGHT TEST

AA1017_2
SA

Figure 7A-23. Cruise - Pressure Altitude 14,000 Feet (Sheet 2 of 5)
7A-101

TM 1-1520-237-10

CRUISE

CRUISE
14,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 14,000 FT
0o C

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170

TRUE AIRSPEED ~ KTS

140
MAX
RANGE

130

150

140

120

130

180

120

170
160
150

110
MAX
RANGE
100

140
130
120

110

110
80

100

100
70

TRUE AIRSPEED ~ KTS

150

LE ~ 30 MINUTES

160

TORQUE AVAILAB

170

~ CONTINUOUS

180

ATF = 0.9
ATF = 1.0

TORQUE AVAILABLE
~ 30 MINUTES

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

ATF = 0.9
ATF = 1.0

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
60

MAX END
AND R / C

MAX END
AND R / C
50

70
22

60
12

70
22

60
12

GW ~
1000 LB

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 75%

PLACARD TORQUE LIMIT = 73%

DATA BASE: FLIGHT TEST

Figure 7A-23. Cruise - Pressure Altitude 14,000 Feet (Sheet 3 of 5)
7A-102

AA1017_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
14,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 14,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160
ATF = 0.9
ATF = 1.0

140
MAX
RANGE
130
120

110
MAX
RANGE

100

110

ATF = 1.0

ATF = 0.9

120

160
150
140
130
120
110

100

70
90

TRUE AIRSPEED ~ KTS

150

170

TES

130

180

TORQUE AVAILABLE ~ 30 MINU

160

TRUE AIRSPEED ~ KTS

140

TES

170

150

TORQUE AVAILABLE ~ 30 MINU

180

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
60
MAX END
AND R / C

MAX END
AND R / C

50
70

70
40

22
12

GW ~
1000 LB

20
30

0
10

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 70%

PLACARD TORQUE LIMIT = 67%

DATA BASIS: FLIGHT TEST

AA1017_4A
SA

Figure 7A-23. Cruise - Pressure Altitude 14,000 Feet (Sheet 4 of 5)
7A-103

TM 1-1520-237-10

CRUISE

CRUISE
14,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 14,000 FT

30oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

10
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

TRUE AIRSPEED ~ KTS

150
MAX
RANGE

140

ATF = 1.0

130

130
120

120

110

100

110
100

INDICATED AIRSPEED ~ KTS

170

140

TORQUE AVAILABLE ~ 30 MINU
TES

~ CONTINUOUS

180

ATF = 0.9

150

60
MAX END
AND R / C

50
70
40
60
12
50

GW ~
1000 LB

30
20

10
0

0
0

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 65%

DATA BASE: FLIGHT TEST

Figure 7A-23. Cruise - Pressure Altitude 14,000 Feet (Sheet 5 of 5)
7A-104

AA1017_5
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T701C (2)

PRESS ALT: 14000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160
10

150

180

120
MAX
RANGE

100
MAX
RANGE
90

TORQUE AVAI

110

LABLE ~ 30 MI

130

100
90

150
140

LABLE ~ 30 MI

140

120

130
120
110

TORQUE AVAI

150

160

NUTES AND CO

130

170

100
90

60
MAX END
AND R / C

MAX END
AND R / C

40
50

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

140

NUTES AND CO

160

NTINUOUS

180

NTINUOUS

170

TRUE AIRSPEED ~ KTS

GW ~
1000 LB

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 90%

PLACARD TORQUE LIMIT = 86%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1018_1
SA

Figure 7A-24. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 1 of 5)
7A-105

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T701C (2)

PRESS ALT: 14000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

LE ~ 30 MINUTES

130
MAX
RANGE

120

110

100
MAX
RANGE
90

TORQUE AVAILAB

TRUE AIRSPEED ~ KTS

140

120

110
100
90

50
MAX END
AND R / C

50
14

130
120
110
100

70
MAX END
AND R / C

22
12

GW ~
1000 LB

50
40

20
30

140

22
12

150

90
60

160

170

TORQUE AVAILABLE
~ 30

150

130

TRUE AIRSPEED ~ KTS

S
AND CONTINUOU

160

180

~ CONTINUOUS

140

170

MINUTES

180

30
20

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 82%

PLACARD TORQUE LIMIT = 79%

0
100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-24. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 2 of 5)
7A-106

AA1018_2
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T701C (2)

PRESS ALT: 14000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
5

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

ATF = 0.9

140

180

130

140
130
MAX
RANGE

120

110

TORQUE AVAILABLE ~ 30

150

110

TORQUE AVAILABLE ~ 30
MINUTES

MINUTES

160

170

100
MAX
RANGE
90

160
150
140
130
120
110

100

100
70

TRUE AIRSPEED ~ KTS

170

TRUE AIRSPEED ~ KTS

150

ATF = 0.9
ATF = 1.0

180

ATF = 1.0

160

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

90
60

80
50

MAX END
AND R / C

60
22
30

50
40

GW ~
1000 LB

30
20

10
0

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 75%

PLACARD TORQUE LIMIT = 73%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1018_3
SA

Figure 7A-24. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 3 of 5)
7A-107

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T701C (2)

PRESS ALT: 14000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

130

150
140
MAX
RANGE

120

110

100
MAX
RANGE

180
170
160
150
140
130
120

110

100

110

80
MAX END
AND R / C

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

140

TORQUE AVAILABLE ~ 30 MINU
TES

160

TRUE AIRSPEED ~ KTS

30
~ CONTINUOUS

ATF = 0.9

170

130

150

TORQUE AVAILABLE ~ 30 MINU
TES

180

ATF = 1.0

160

ATF = 1.0

~ CONTINUOUS

ATF = 0.9

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

70
40

60
12

20
30

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 70%

PLACARD TORQUE LIMIT = 67%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-24. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 4 of 5)
7A-108

AA1018_4
SA

TM 1-1520-237-10

CRUISE

CRUISE
14000 FT
T701C (2)

PRESS ALT: 14000 FT
o

30 C

TOTAL FUEL FLOW ~ 100 LB/HR
4

12
170

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160
20

160

TRUE AIRSPEED ~ KTS

130

150
140
MAX
RANGE

120

110

100

110
70

100

INDICATED AIRSPEED ~ KTS

170

130

140

TORQUE AVAILABLE ~ 30 MINUTES

ATF = 0.9

180

150

ATF = 1.0

~ CONTINUOUS

90
80

MAX END
AND R / C

70
40
60

50
12

18
20

40
GW ~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMIT = 65%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1018_5
SA

Figure 7A-24. Cruise High Drag - Pressure Altitude 14,000 Feet (Sheet 5 of 5)

7A-109

TM 1-1520-237-10

CRUISE

CRUISE
16,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 16,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

180
150

180

170
170
CONTINUOUS

MINUTES AND CONT
INUOUS

160

MAX
RANGE

120

110

100

90
80

MAX
RANGE

100

140
130
120

110

100
90
80

60
MAX END
AND R / C

MAX END
AND R / C

150

~ 30 MINUTES AND

130

TORQUE AVAILABLE
~ 30

TRUE AIRSPEED ~ KTS

140

120

160

TORQUE AVAILABLE

150

130

40
50

GW ~
1000 LB

30
20

TRUE AIRSPEED ~ KTS

140

30
20

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 82%

PLACARD TORQUE LIMIT = 79%

DATA BASE: FLIGHT TEST

Figure 7A-25. Cruise - Pressure Altitude 16,000 Feet (Sheet 1 of 4)
7A-110

AA1019_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
16,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 16,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150
180

130

MAX
RANGE

120

110

100
90

MAX
RANGE

100

TORQUE AVAILABLE

120

40
16

110

MAX END
AND R / C

80
70
60

GW ~
1000 LB

120

130

100

140

MAX END
AND R / C

150

160

LE ~ 30 MINUTES

140

170

TORQUE AVAILAB

150

180

130

~ 30 MINUTES AND CO

160

TRUE AIRSPEED ~ KTS

140

NTINUOUS

170

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

GW ~
1000 LB

30
20

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT =76%

PLACARD TORQUE LIMIT = 72%

DATA BASE: FLIGHT TEST

AA1019_2
SA

Figure 7A-25. Cruise - Pressure Altitude 16,000 Feet (Sheet 2 of 4)
7A-111

TM 1-1520-237-10

CRUISE

CRUISE
16,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 16,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

MAX
RANGE

120

TES

170

120

110
MAX
RANGE

100

110

180

160

150
140
130
120
110

100

70
90

TRUE AIRSPEED ~ KTS

140
130

130

TORQUE AVAILABLE ~ 30

TRUE AIRSPEED ~ KTS

150

140

ATF = 0.9

ATF = 0.9
ATF = 1.0

160

TORQUE AVAILABLE ~ 30 MINU

170

150

MINUTES

180

ATF = 1.0

160

~ CONTINUOUS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

90
60
MAX END
AND R / C

MAX END
AND R / C

50
70

70
40

60
12

50
GW ~
1000 LB

GW ~
1000 LB

20
30

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 69%

PLACARD TORQUE LIMIT = 66%

DATA BASE: FLIGHT TEST

Figure 7A-25. Cruise - Pressure Altitude 16,000 Feet (Sheet 3 of 4)
7A-112

AA1019_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
16,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 16,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

MAX
RANGE

130
120
110

ATF = 0.9

ATF = 1.0

110
MAX
RANGE

100

100
90
MAX END
AND R / C

170
160
150
140
130
120
110

100

90
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

120

180

TORQUE AVAILABLE ~ 30 MINUTES

TRUE AIRSPEED ~ KTS

150
140

130

TORQUE AVAILABLE ~ 30 MINU

160

~ CONTINUOUS

170

140

TES

180

ATF = 0.9
ATF = 1.0

~ CONTINUOUS

150

70
40

GW ~
1000 LB

50
40

10
0

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 64%

PLACARD TORQUE LIMIT = 61%

DATA BASE: FLIGHT TEST

AA1019_4
SA

Figure 7A-25. Cruise - Pressure Altitude 16,000 Feet (Sheet 4 of 4)
7A-113

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T701C (2)

PRESS ALT: 16000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

140
130

110
100

110

100
MAX
RANGE
90

60
MAX END
AND R / C

NTINUOUS

120

TORQUE AVAIL

MAX
RANGE

120

130

ABLE ~ 30 MINU

150

180

160
150
140
130
120
110
100
90
80

MAX END
AND R / C

170

~ 30 MINUTES AND CO

160

TRUE AIRSPEED ~ KTS

140

TES AND CONT

170

TORQUE AVAILABLE

180

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

INUOUS

150

30
20

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 83%

PLACARD TORQUE LIMIT = 79%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-26. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 1 of 4)
7A-114

AA1020_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T701C (2)

PRESS ALT: 16000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

130
MAX
RANGE

120

110

100

110
100
90

MAX
RANGE

170
160
150

TORQUE AVAILABLE

140

120

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

150

180

140
130
120
110

100

TRUE AIRSPEED ~ KTS

~ 30 MINUTES AND CO

160

130

~ 30 MINUTES

NTINUOUS

170

~ CONTINUOUS

140
180

90
60

80
50
MAX END
AND R / C

MAX END
AND R / C

60
12

50
40

GW ~
1000 LB

50
40

GW ~
1000 LB

30
20

0
20

100

0
20

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 75%

PLACARD TORQUE LIMIT = 72%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1020_2
SA

Figure 7A-26. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 2 of 4)
7A-115

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T701C (2)

PRESS ALT: 16000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

130

MAX
RANGE

120
110

ATF = 0.9
ATF = 1.0

100
MAX
RANGE

100
90

160
150
140
130
120
110

100

50
70

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

140

110

170

TES

ATF = 0.9
ATF = 1.0

150

120

180

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

160

130

~ CONTINUOUS

170

TES

180

140

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

150

MAX END
AND R / C
70

60
20

30
50

50
12

30
20

GW ~
1000 LB

0
20

100

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 69%

PLACARD TORQUE LIMIT = 66%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-26. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 3 of 4)
7A-116

AA1020_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
16000 FT
T701C (2)

PRESS ALT: 16000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

140
MAX
RANGE

130
120
110

ATF = 0.9

TES

100
MAX
RANGE

ATF = 1.0

110

170

160
150
140
130
120
110

100

TRUE AIRSPEED ~ KTS

150

120

180

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

160

~ CONTINUOUS

170

130

TES

~ CONTINUOUS

180

140

TORQUE AVAILABLE ~ 30 MINU

ATF = 0.9

150

100
60

90
80

MAX END
AND R / C

GW ~
1000 LB

50
40

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMITS = 64%

0
10

TORQUE PER ENGINE ~ %
PLACARD TORQUE LIMITS = 61%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1020_4
SA

Figure 7A-26. Cruise High Drag - Pressure Altitude 16,000 Feet (Sheet 4 of 4)
7A-117

TM 1-1520-237-10

CRUISE

CRUISE
18,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 18,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150
180

180

MAX
RANGE

S
110

TORQUE AVAILAB

100

110

100

MAX
RANGE

150
140
130
120
110
100

TRUE AIRSPEED ~ KTS

130

AND CONTINUOU

120

160

LE ~ 30 MINUTES

140

120

130

LE ~ 30 MINUTES

150

170

TORQUE AVAILAB

160

TRUE AIRSPEED ~ KTS

140

AND CONTINUOU

170

80
70

40
12

70
60

12
30

GW ~
1000 LB

MAX END
AND R / C

GW ~
1000 LB

10
0

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 76%

PLACARD TORQUE LIMIT = 72%

DATA BASE: FLIGHT TEST

Figure 7A-27. Cruise - Pressure Altitude 18,000 Feet (Sheet 1 of 4)
7A-118

AA1021_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
18,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 18,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

180

140

30 MINUTES & CONTI
NUOUS

140
130
120
110
100
90

160

120

150
110
140
100

TORQUE AVAILABLE~

MAX
RANGE

60
80

170

MAX END
AND R / C

130
120
110
100

TRUE AIRSPEED ~ KTS

150

130

MINUTES

160

180

TORQUE AVAILABLE ~ 30

170

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

150

90
80

70
40

50
12

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 69%

PLACARD TORQUE LIMIT = 66%

DATA BASE: FLIGHT TEST

AA1021_2
SA

Figure 7A-27. Cruise - Pressure Altitude 18,000 Feet (Sheet 2 of 4)
7A-119

TM 1-1520-237-10

CRUISE

CRUISE
18,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 18,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

110

ATF = 0.9

ATF = 1.0

160
150

MAX
RANGE

140

130
120
110

100
90

100

60
MAX END
AND R / C

TRUE AIRSPEED ~ KTS

120

100

170

TORQUE AVAILABLE

MAX
RANGE

130

110

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

150

120

180

~ 30 MINUTES

ATF = 0.9

130

160

140

~ CONTINUOUS

170

ATF = 1.0

180

140

~ 30 MINUTES

~ CONTINUOUS

150

90
MAX END
AND R / C

70
40

60
12

18
50

GW ~
1000 LB

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 63%

PLACARD TORQUE LIMIT = 61%

DATA BASE: FLIGHT TEST

Figure 7A-27. Cruise - Pressure Altitude 18,000 Feet (Sheet 3 of 4)
7A-120

AA1021_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
18,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 18,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
10

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
160

150

120
110

TES ATF = 1.0

170

100
MAX
RANGE
90

160
150
140
130
120
110

70
100

TRUE AIRSPEED ~ KTS

130

180

TORQUE AVAILABLE ~ 30 MINU

MAX
RANGE

110

ATF = 0.9

ATF = 1.0

ATF = 0.9

150
140

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

160

120

TES

170

130

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

180

140

100
60

90
MAX END
AND R / C

MAX END
AND R / C

40
60
GW ~ 12
1000 LB

GW ~ 12
1000 LB

50
40

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 59%

PLACARD TORQUE LIMIT = 56%

DATA BASE: FLIGHT TEST

AA1021_4
SA

Figure 7A-27. Cruise - Pressure Altitude 18,000 Feet (Sheet 4 of 4)
7A-121

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T701C (2)

PRESS ALT: 18000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
13

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160
10

160
150
140

120

MAX
RANGE

110
100

NTINUOUS

120

110

100

TORQUE AVAILABLE

130

180
170

~ 30 MINUTES AND CO

130

160
150
140
130

TORQUE AVAILABLE

170

TRUE AIRSPEED ~ KTS

140

~ 30 MINUTES AND CO

180

MAX
RANGE

120
110
100

TRUE AIRSPEED ~ KTS

NTINUOUS

150

90
60

80
MAX END
AND R / C

MAX END
AND R / C

60
12

50
GW ~
1000 LB

20
50

GW ~
1000 LB

40
30

10
0

0
20

100

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 76%

PLACARD TORQUE LIMIT = 72%

100

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-28. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 1 of 4)
7A-122

AA1022_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T701C (2)

PRESS ALT: 18000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
20

160

30
~ CONTINUOUS

150

AND CONTINUOU

140

140
130
MAX
RANGE

120

110

LE ~ 30 MINUTES

150

110
100
90

180

120

100

TORQUE AVAILAB

TRUE AIRSPEED ~ KTS

160

130

LE ~ 30 MINUTES

170

170
160
150

TORQUE AVAILAB

180

MAX
RANGE

140
130
120
110

100
90

80
70

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

MAX END
AND R / C
70

40
60

60
30

50
12

18
40

20
GW ~
1000 LB

GW ~
1000 LB

30
20

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 69%

PLACARD TORQUE LIMIT = 66%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1022_2
SA

Figure 7A-28. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 2 of 4)
7A-123

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T701C (2)

PRESS ALT: 18000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS
11

TOTAL FUEL FLOW ~ 100 LB/HR
4

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

ATF = 0.9
ATF = 1.0

180

120

150
140
130

MAX
RANGE

120
110
100

110

100

170

TORQUE AVAILABLE ~ 30
MINUTES

160

MAX
RANGE

160
150
140
130
120
110
100

90
80

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

130

170

TRUE AIRSPEED ~ KTS

~ CONTINUOUS

180

140

ATF = 0.9
ATF = 1.0

~ CONTINUOUS

150

30
GW ~
1000 LB

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 63%

PLACARD TORQUE LIMIT = 61%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-28. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 3 of 4)
7A-124

A1022_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
18000 FT
T701C (2)

PRESS ALT: 18000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
5

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

MAX
RANGE

120
110
100

100

90
MAX
RANGE

160

TORQUE AVAILABLE ~ 30 MINU

140

170

TES

150

130

110

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

160

180

150
140
130
120
110

TRUE AIRSPEED ~ KTS

120

170

ATF = 0.9

130

ATF = 1.0

~ CONTINUOUS

180

140

ATF = 0.9
ATF = 1.0

~ CONTINUOUS

150

100
60

MAX END
AND R / C

70
60
50

GW ~
1000 LB

20
10

10
0

0
10

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 59%

PLACARD TORQUE LIMIT = 56%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1022_4
SA

Figure 7A-28. Cruise High Drag - Pressure Altitude 18,000 Feet (Sheet 4 of 4)
7A-125

TM 1-1520-237-10

CRUISE

CRUISE
20,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 20,000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

180

140

180

170
130

170

110
100
90
80

100

MAX END
AND R / C

MAX
RANGE

150
140
130
120

60
MAX END
AND R / C

110
100

TRUE AIRSPEED ~ KTS

MAX
RANGE

110

TORQUE AVAILABLE~

130

160

30 MINUTES & CONTI

140

120

TORQUE AVAILABLE~

TRUE AIRSPEED ~ KTS

150

NUOUS

160

90
80

70
40

60
GW ~
1000 LB

40
20

10
0

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 69%

PLACARD TORQUE LIMIT = 66%

DATA BASE: FLIGHT TEST

Figure 7A-29. Cruise - Pressure Altitude 20,000 Feet (Sheet 1 of 4)
7A-126

AA1023_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
20,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 20,000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
2

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

140

160

140
MAX
RANGE

130
120
110
100
90
MAX END
AND R / C

110

100

TORQUE AVAILABLE~30 MINU

TRUE AIRSPEED ~ KTS

150

120

MAX
RANGE

170
160
150
140
130
120
110

TRUE AIRSPEED ~ KTS

TES & CONTINUOUS

170

180

TES

130

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

180

100

90
MAX END
AND R / C

70
40

60
GW ~ 12
1000 LB

GW ~
1000 LB

50
40

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 63%

PLACARD TORQUE LIMIT = 60%

DATA BASE: FLIGHT TEST

AA1023_2
SA

Figure 7A-29. Cruise - Pressure Altitude 20,000 Feet (Sheet 2 of 4)
7A-127

TM 1-1520-237-10

CRUISE

CRUISE
20,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 20,000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

140
MAX
RANGE

130
120
110

100
MAX
RANGE

TES ATF = 1.0

ATF = 0.9

110

170

160
150
140
130
120
110

100

TRUE AIRSPEED ~ KTS

150

120

180

TORQUE AVAILABLE ~ 30 MINU

TRUE AIRSPEED ~ KTS

160

~ CONTINUOUS

170

130

TORQUE AVAILABLE ~ 30 MINU

~ CONTINUOUS

180

140

TES ATF = 1.0

ATF = 0.9

150

100
60

90
MAX END
AND R / C

MAX END
AND R / C

80
70

GW ~ 12
1000 LB

50
40

10
0

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 58%

PLACARD TORQUE LIMIT = 56%

DATA BASE: FLIGHT TEST

Figure 7A-29. Cruise - Pressure Altitude 20,000 Feet (Sheet 3 of 4)
7A-128

AA1023_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
20,000 FT
T701C (2)

CLEAN CONFIGURATION
PRESS ALT: 20,000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
4

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT OF F
10

160

150

140
MAX
RANGE

130
120

110

ATF = 0.9

100

90
MAX
RANGE
80

110

180
170
160
150
140
130
120
110

100

60
90

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

150

120

TORQUE AVAILABLE ~ 30 MINUTES ATF = 1.0

160

130

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

170

ATF = 0.9

180

TORQUE AVAILABLE ~ 30 MINUTES ATF = 1.0

~ CONTINUOUS

140

MAX END
AND R / C

GW ~ 12
1000 LB

70
GW ~ 12
1000 LB

40
30

30
20

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT =53%

PLACARD TORQUE LIMIT = 51%

DATA BASE: FLIGHT TEST

AA1023_4
SA

Figure 7A-29. Cruise - Pressure Altitude 20,000 Feet (Sheet 4 of 4)
7A-129

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T701C (2)

PRESS ALT: 20000 FT
−40oC

−50oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

140

130
NTINUOUS

180

170

TORQUE AVAILABLE ~ 30
MINUTES AND CONTINU
OUS

140
130
120
MAX
RANGE

110
100
90
80
MAX END
AND R / C

110

100

90
MAX
RANGE

160
150
140
130
120

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

150

170

~ 30 MINUTES AND CO

120

160

110
100
90
80

MAX END
AND R / C
70

60
12

TRUE AIRSPEED ~ KTS

180

50
GW ~
1000 LB

GW ~
1000 LB

10
0

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 69%

PLACARD TORQUE LIMIT = 66%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-30. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 1 of 4)
7A-130

AA1024_1
SA

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T701C (2)

PRESS ALT: 20000 FT
−20oC

−30oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS

TOTAL FUEL FLOW ~ 100 LB/HR

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

150

140

130

130
MAX
RANGE

110
100
90
80

100

90
MAX
RANGE
80

MAX END
AND R / C

180

TES

170
160
150
140
130
120
110
100

60
90
50

MAX END
AND R / C

30
50

GW ~
1000 LB

TRUE AIRSPEED ~ KTS

140

120

110

~ 30 MINUTES AND CO

150

TORQUE AVAILABLE

TRUE AIRSPEED ~ KTS

160

120

TORQUE AVAILABLE ~ 30 MINU

170

~ CONTINUOUS

NTINUOUS

180

12
20

50
40

GW ~
1000 LB

0
10

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 63%

PLACARD TORQUE LIMIT = 60%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1024_2
SA

Figure 7A-30. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 2 of 4)
7A-131

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T701C (2)

PRESS ALT: 20000 FT
0 oC

−10oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
10

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

160

150

130
MAX
RANGE

110

ATF = 0.9
ATF = 1.0

100

90
MAX
RANGE

TES

140

120

110

180
170
160
150
140
130
120
110

100

TRUE AIRSPEED ~ KTS

150

~ CONTINUOUS

160

TORQUE AVAILABLE ~ 30 MINU

170

TRUE AIRSPEED ~ KTS

120

~ CONTINUOUS

180

130

TORQUE AVAILABLE ~ 30 MINU

ATF = 0.9
ATF = 1.0

140

100

60
90

MAX END
AND R / C

70
60
12

50
40

GW ~
1000 LB

40
30

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 58%

PLACARD TORQUE LIMIT = 56%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

Figure 7A-30. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 3 of 4)
7A-132

AA1024_3
SA

TM 1-1520-237-10

CRUISE

CRUISE
20000 FT
T701C (2)

PRESS ALT: 20000 FT
20oC

10oC

TOTAL FUEL FLOW ~ 100 LB/HR
3

IAS ~ KTS
10

TOTAL FUEL FLOW ~ 100 LB/HR
3

170
TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
10

TRQ ~ % FOR DRAG
AREA OF 10 SQ FT
160

150

130
MAX
RANGE

120
110
100

ATF = 0.9
ATF = 1.0

ATF = 1.0

MAX
RANGE

170
160
150
140
130
120
110

100

90
80

MAX END
AND R / C

TRUE AIRSPEED ~ KTS

140

100

TES

150

110

180

TORQUE AVAILABLE ~ 30 MINU

160

~ CONTINUOUS

TRUE AIRSPEED ~ KTS

170

120

TES

~ CONTINUOUS

180

130

TORQUE AVAILABLE ~ 30 MINU

ATF = 0.9

140

MAX END
AND R / C

70
12

60
50

GW ~
1000 LB

60
50

GW ~
1000 LB

30
10

0
0

TORQUE PER ENGINE ~ %

PLACARD TORQUE LIMIT = 53%

PLACARD TORQUE LIMIT = 51%

NOTE: SHORT DASH LINES: FERRY MISSION ONLY
DATA BASE: FLIGHT TEST

AA1024_4
SA

Figure 7A-30. Cruise High Drag - Pressure Altitude 20,000 Feet (Sheet 4 of 4)
7A-133

TM 1-1520-237-10

Section V OPTIMUM CRUISE
7A.20 OPTIMUM RANGE CHARTS.
This section presents a method to optimize cruise performance for long range missions when the altitudes flown
are not restricted by other requirements. The optimum altitude for maximum range charts (Figures 7A-31 and 7A-32)
provides the pressure altitude at which to cruise to obtain
the maximum possible range for any gross weight and FAT
conditions. The altitude determined for optimum range may

7A-134

also be used for optimum endurance. Enter the chart at a
current cruise or takeoff temperature condition and move
along the temperature guide lines to the anticipated gross
weight for cruise and obtain the optimum pressure altitude.
Turn to the cruise chart closest to the altitude and temperature predicted by the optimum range chart for specific
cruise information. The use of this chart is shown by the
example.

TM 1-1520-237-10

OPTIMUM RANGE
CLEAN CONFIGURATION 100% RPM R
HIRSS (BAFFLES INSTALLED)

GROSS WEIGHT ~ 1000 LB
22

EXAMPLE

WANTED
CRUISE ALTITUDE FOR OPTIMUM RANGE
AND CORRESPONDING CRUISE CHART FOR
FLIGHT CONDITIONS

KNOWN

METHOD
ENTER CHART AT FAT (24oC). MOVE
RIGHT TO REFERENCE / OPTIMUM
PRESSURE ALTITUDE (1,500 FT). MOVE
PARALLEL WITH THE TEMPERATURE
TREND LINE TO AIRCRAFT GROSS
WEIGHT (16,600 LB). MOVE LEFT
OR RIGHT PARALLELING THE
TEMPERATURE TREND LINE TO
THE NEAREST EVEN THOUSAND
REFERENCE / OPTIMUM PRESSURE
ALTITUDE LINE (12,000). MOVE
LEFT TO FREE AIR TEMPERATURE
LINE (30). MOVE UP OR DOWN
TO NEAREST TEN VALUE ON THE
FREE AIR TEMPERATURE SCALE (0oC).
SELECT CRUISE CHART WITH
ALTITUDE AND TEMPERATURE
DATA AT THE NEAREST REFERENCE /
OPTIMUM PRESSURE ALTITUDE
(12,000 FT) AND THE NEAREST
TEN DEGREE FREE AIR TEMPERATURE
(0oC).

FREE AIR TEMPERATURE ~ OC

REFERENCE CONDITIONS OF:
PRESSURE ALTITUDE = 1,500 FT
FAT = 24OC
GW = 16,600 LB

−10

−20

TEM
PER
ATU
TRE
RE
ND
LIN
ES

−30

−40

−50

−60
0

OPTIMUM PRESSURE ALTITUDE ~ 1000 FT

DATA BASIS:

FLIGHT TEST

AA1256B
SA

Figure 7A-31. Optimum Altitude For Maximum Range

7A-135

TM 1-1520-237-10

OPTIMUM RANGE
HIGH DRAG CONFIGURATION 100% RPM R
HIRSS (BAFFLES INSTALLED)

GROSS WEIGHT ~ 1000 LBS
24 23 22

FREE AIR TEMPERATURE ~ DEG C

20
TEM

PER

ATU

TRE

LIN

−10

−20

−30

NLY
RY O

−40

FER

−50

−60
0

OPTIMUM PRESSURE ALTITUDE ~ 1000 FT

AA1026A
SA

Figure 7A-32. Optimum Altitude For Maximum Range - High Drag
7A-136

TM 1-1520-237-10

Section VI DRAG
7A.21 EXTERNAL LOAD DRAG CHART.
The general shapes of typical external loads are shown
on Figure 7A-33. as a function of the load frontal area. The
frontal area is combined with the typical drag coefficient of
the general shapes to obtain a drag multiplying factor for
use with the 10 sq. ft. drag scale on each cruise chart. The
TRQ ˜% value obtained from the cruise chart is multiplied by the drag multiplying factor and added to indicated
torque to obtain total torque required at any airspeed.

7A.22 AIRCRAFT
CONFIGURATION
CHANGES FOR USE WITH CLEAN
CHARTS.

DRAG
CRUISE

Table 7A-1. Configuration Drag Change
DRAG CHANGES FOR USE WITH CLEAN CRUISE CHARTS
Change in Flat Plate Drag
F Sq. Ft.
Area -

Drag Multiplying Factor

a. Both cargo doors open

6.0

0.60

b. Cargo doors removed

4.0

0.40

c. Cargo mirror installed

0.3

0.03

d. IR Countermeasure Transmitter (ALQ-144) installed

0.8

0.08

e. Chaff Dispenser installed

0.3

0.03

f. Blade Erosion Kit

2.0

0.20

g. Skis installed

3.0

0.30

Item

7A.23 AIRCRAFT
CONFIGURATION
DRAG
CHANGES FOR USE WITH HIGH DRAG CRUISE
CHARTS.

ternal drag load method. Typical high drag configuration
changes that have been established from flight test or analysis along with the drag multiplying factors are shown.

When external equipment differs from the baseline high
drag configuration as defined in this Section, a drag correction should be made using Figure 7A-34 similar to the ex-

7A-137

TM 1-1520-237-10

EXTERNAL DRAG LOAD METHOD:

EXAMPLE

KNOWN:

WANTED:

ENTER CHART AT SYMBOL
FOR CYLINDER
MOVE RIGHT TO 80 SQ FT.
MOVE DOWN AND READ
DRAG MULTIPLYING FACTOR
= 4.5

SHAPE OF EXTERNAL
LOAD = CYLINDER
FRONTAL AREA OF
EXTERNAL LOAD = 80 SQ FT

DRAG MULTIPLYING FACTOR
DUE TO EXTERNAL LOAD

LOAD
DRAG

INCREASE IN DRAG AREA DUE TO EXTERNAL LOAD
0

100

120

140

160

180

200

220

240

SPHERE

STREAMLINED
CYLINDER

CYLINDER
CUBE
FLAT
PLATE
BOX

100

FRONTAL AREA
OF EXTERNAL
LOAD ~ SQ FT

BOX
IN
NET
0

DATA BASIS:

ESTIMATED

Figure 7A-33. External Load Drag

7A-138

DRAG MULTIPLYING FACTOR

24
AA0684A
SA

TM 1-1520-237-10

DRAG CONFIGURATIONS

CHANGE
IN
FLAT
PLATE
DRAG
F
SQ FT

DRAG
MULTI−
PLYING
FACTOR

ESSS − CLEAN, PYLONS REMOVED

−4.0

−0.40

ESSS − FOUR PYLONS / NO STORES

−1.7

−0.17

ESSS−TWO 450−GALLON TANKS INBOARD

0.5

0.05

−TWO 230−GALLON TANKS INBOARD

0.0

0.00

2.5

0.25

2.0

0.20

32.5

3.25

10.5

1.05

SKIS INSTALLED

3.0

0.30

BOTH CARGO DOORS OPEN

6.0

0.60

BOTH CARGO DOORS REMOVED

4.0

0.40

CARGO MIRROR INSTALLED

0.3

0.03

IR COUNTERMEASURE TRANSMITTER (ALQ−144) REMOVED

−0.8

−0.08

CHAFF DISPENSER REMOVED

−0.3

−0.03

HIGH DRAG
CRUISE CHART BASELINE
SPECIAL MISSION EQUIPMENT CONFIGURATIONS

ESSS−TWO 230−GALLON TANKS OUTBOARD
−TWO 450−GALLON −TANKS INBOARD

ESSS − FOUR 230−GALLON TANKS

VOLCANO SYSTEM INSTALLED (BOTH RACKS)
* VOLCANO CORRECTION MUST BE USED WITH HIGH DRAG CHARTS ONLY
VOLCANO SYSTEM INSTALLED (LOWER RACKS ONLY)

AA0685B
SA

Figure 7A-34. Typical High Drag Configurations
Change 8

7A-139

TM 1-1520-237-10

Section VII CLIMB-DESCENT
7A.24 CLIMB/DESCENT CHART.
The CLIMB/DESCENT chart (Figures 7A-35 and 7A36) presents the rate of climb or descent resulting from an
increase or decrease of engine torque from the value required for level flight above 40 KIAS. The data are presented at 100% RPM R for various gross weights. The
charts may also be used in reverse to obtain the torque

7A-140

increase or reduction required to achieve a desired steady
rate of climb or descent. The maximum R/C may be determined by subtracting the cruise chart torque required from
the maximum torque available at the desired flight conditions. Then enter the difference on the torque increase scale
of the climb chart, move up to the gross weight, and read
the resulting maximum R/C.

TM 1-1520-237-10

CLIMB/DESCENT
100% RPM R

CLIMB/DESCENT

CLEAN CONFIGURATION
FOR AIRSPEED ABOVE 40 KIAS
12

4000

DESCENT

3500

GROSS
WEIGHT
~ 1000 LB

RATE OF DESCENT ~ FT. / MIN.

20
22

3000

2500

2000

1500

1000

500

0
0

TORQUE REDUCTION − PER ENGINE ~% TRQ
12

EXAMPLE
ENTER AT 1000 FT / MIN
MOVE RIGHT TO 20,000 LB
MOVE DOWN READ TRQ
INCREASE = 28%

RATE OF CLIMB ~ FT. / MIN.

3500

3000

GROSS
WEIGHT
~ 1000 LB

CLIMB

20
22

2500

2000

1500

1000

500

0
0

TORQUE INCREASE − PER ENGINE ~% TRQ
AA1638

DATA BASIS: FLIGHT TEST

Figure 7A-35. Climb/Descent
7A-141

TM 1-1520-237-10

CLIMB DESCENT

CLIMB /
DESCENT

100% RPM R
FOR AIRSPEED ABOVE 40 KIAS
4000
GROSS
WEIGHT
~ 1000 LB
3500

16
18

DESCENT

RATE OF DESCENT ~ FT/MIN

20
3000

22
24

2500

2000

1500

1000

500

0
0

TORQUE REDUCTION − PER ENGINE ~ % Q
3500

12
CLIMB

3000

GROSS
WEIGHT
~ 1000 LB

RATE OF CLIMB ~ FT/MIN

18
20

2500

22
24
2000

1500

1000

500

0
0

TORQUE INCREASE − PER ENGINE ~ % Q
DATA BASIS:

FLIGHT TEST

AA1027
SA

Figure 7A-36. Climb/Descent - High Drag

7A-142

TM 1-1520-237-10

Section VIII FUEL FLOW
7A.25 IDLE FUEL FLOW.
Dual-engine idle fuel flow is presented as a function of
altitude at 0°C FAT in Table 7A-2.

c. When the IR suppressor system is operating in the
benign mode (exhaust baffles removed), the fuel flow will
decrease about 7 lbs/hr.
7A.27 DUAL-ENGINE FUEL FLOW.

7A.26 SINGLE-ENGINE FUEL FLOW.
Engine fuel flow is presented in Figure 7A-37 for various torque and pressure altitudes at a baseline FAT of 0°C
with engine bleed air extraction off. When operating at
other than 0°C FAT, engine fuel flow is increased 1% for
each 20°C above the baseline temperature and, decreased
1% for each 20°C below the baseline temperature.
To determine single-engine fuel flow during cruise, enter the fuel flow chart at double the torque required for
dual-engine cruise as determined from the cruise charts and
obtain fuel flow from the single-engine scale. The singleengine torque may not exceed the transmission limit shown
on the chart. With bleed air extracted, fuel flow increases as
follows:

Dual-engine fuel flow may be obtained from Figure
7A-37 when each engine is indicating approximately the
same torque by averaging the indicated torques and reading
fuel flow from the dual-engine fuel flow scale. When operating at other than the 0° FAT baseline, dual-engine fuel
flow is increased 1% for each 20°C above baseline and is
decreased 1% for each 20°C below baseline temperature.
With bleed air extracted, dual engine fuel flow increases as
follows:
a. Engine anti-ice on - +100 lbs/hr.
b. Cockpit heater on - +12 lbs/hr.
c. When the IR suppressor system is operating in the
benign mode (exhaust baffles removed), the fuel flow decreases about 14 lbs/hr.

a. Engine anti-ice on - +50 lbs/hr.
b. Cockpit heater on - +6 lbs/hr.

Table 7A-2. Idle Fuel Flow
Pressure Altitude Feet

Ng = 63-71%
Ground Idle (66% RPM R)
Lb/Hr

Ng = 79-88%
Flat Pitch (100% RPM R)
Lb/Hr

APU (Nominal) Lb/Hr

360

590

120

4,000

310

515

105

8,000

280

445

12,000

250

400

16,000

220

340

20,000

185

280

7A-143

TM 1-1520-237-10

SINGLE/DUAL−ENGINE FUEL FLOW
HIRSS SUPPRESSED MODE BLEED AIR OFF
100% RPM R FAT= 0 o C

DUAL−ENGINE FUEL FLOW ~ LB/HR
200

400

600

800

1000

1200

1400

1600

1800

2000

900

1000

140
TRANSMISSION LIMIT − 1 ENGINE

130
NOTE
INCREASE FUEL FLOW 1%
FOR EACH 20o C ABOVE
0o C FAT AND DECREASE
FUEL FLOW 1% FOR EACH
20o C BELOW 0o C FAT

120

100

INDICATED TORQUE PER ENGINE ~ %

110

20
20
16

PRESSURE
ALTITUDE ~ 1000 FT

10
100

200

300

400

500

600

700

800

SINGLE−ENGINE FLOW ~ LB/HR

DATA BASIS:
ENGINE MANUFACTURER
SPEC.

AA1028B
SA

Figure 7A-37. Single/Dual-Engine Fuel Flow

7A-144

Change 10

TM 1-1520-237-10

Section IX AIRSPEED SYSTEM CHARACTERISTICS
7A.28 AIRSPEED SYSTEM DESCRIPTION.
NOTE
Indicated airspeeds below 40 KIAS are unreliable. Airspeed conversion data KIAS to
KTAS for speeds above 40 KIAS are provided in Section IV CRUISE.
The UH-60L is equipped with the pitot-static system
with wedge mounted pitot-static probes.
7A.29 AIRSPEED CHARTS.
7A.29.1 Airspeed Correction Charts. All indicated
airspeeds shown on the cruise charts are based on level
flight. Figures 7A-38 and 7A-39 provide the airspeed correction to be added to the cruise chart IAS value to determine the related airspeed indicator reading for other than
level flight mode. There are relatively large variations in
airspeed system error associated with climbs and descents.
These errors are significant and Figures 7A-38 and 7A-39
are provided primarily to show the general magnitude and
direction of the errors associated with the various flight
modes. If desired these figures may be used in the manner
shown by the examples to calculate specific airspeed corrections.

7A.29.2 Airspeed System Dynamic Characteristics. The dynamic characteristics of the pilot and copilot
airspeed indicating systems are normally satisfactory. However, the following anomalies in the airspeed and IVSI indicating system may be observed during the following maneuvers or conditions:
a. During takeoffs, in the speed range of 40 to 80 KIAS,
5 to 10 KIAS airspeed fluctuation may be observed on the
pilot’s and copilot’s airspeed indicators.
b. Power changes in high power, low airspeed climbs
may cause as much as 30 knot airspeed changes in indicated airspeed. Increase in power causes increase in indicated airspeed, and a decrease in power causes decrease in
indicated airspeed.
c. The pilot and copilot airspeed indicators may be unreliable during high power climbs at low airspeeds (less
than 50 KIAS) with the copilot system reading as much as
30 knots lower than the pilot system.
d. In-flight opening and closing of doors and windows
may cause momentary fluctuations of approximately 300
feet per minute on the vertical speed indicators.

7A-145

TM 1-1520-237-10

AIRSPEED SYSTEM CORRECTION
20

CORRECTION TO ADD ~ KNOTS

R / C GREATER THAN 1400 FT / MIN
15

AUTOROTATION

DIVE
LEVEL
FLIGHT

−5
R / C LESS THAN 1400 FT / MIN
−10

−15
20

100

120

140

160

180

IAS FROM CLEAN CRUISE CHARTS ~ KNOTS

EXAMPLE
WANTED:
INDICATED AIRSPEED TO CLIMB AT
MAXIMUM RATE OF CLIMB.

KNOWN:
70 KIAS MAX END / AND R / C FROM
APPROPRIATE CRUISE CHART FOR
A GIVEN PRESSURE ALTITUDE, FAT,
AND GROSS WEIGHT.

METHOD:
ENTER AT KNOWN IAS FROM
CRUISE CHART, MOVE UP TO R / C GREATER
THAN 1,400 FPM, MOVE LEFT READ CORRECTION
TO ADD TO IAS = +12.5 KTS, RE−ENTER
AT KNOWN IAS FROM CRUISE CHART, MOVE UP
TO R / C LESS THAN 1,400 FPM LINE, MOVE LEFT,
READ CORRECTION TO ADD TO IAS = −4 KTS
CALCULATE IAS FOR MAX R / C WHEN:
FOR R / C GREATER THAN 1,400 FPM, AIRSPEED = 70 KIAS +12.5 KIAS = 82.5 KIAS
FOR R / C LESS THAN 1,400 FPM, AIRSPEED = 70 KIAS −4 KIAS = 66 KIAS

DATA BASIS:

AA0321A

FLIGHT TEST

Figure 7A-38. Airspeed Correction Chart

7A-146

TM 1-1520-237-10

AIRSPEED SYSTEM CORRECTION

CORRECTION TO BE ADDED ~ KNOTS

10
AUTOROTATION
5
LEVEL FLIGHT
0
R / C GREATER THAN 1400 FT / MIN
−5

−10
R / C LESS THAN 1400 FT / MIN
−15

−20
20

100

120

140

160

IAS FROM HIGH DRAG CRUISE CHARTS ~ KNOTS

DATA BASIS: FLIGHT TEST

AA1029A
SA

Figure 7A-39. Airspeed Correction Chart - High Drag

7A-147

TM 1-1520-237-10

Section X SPECIAL MISSION PERFORMANCE
7A.30 SPECIAL MISSION FLIGHT PROFILES.
Figures 7A-40 through 7A-42 show special mission
flight profiles required to obtain near maximum range when
equipped with ESSS in three different tank configurations.
The upper segment of each chart provides the recommended altitude profile along with the IAS and average
TRQ versus distance traveled. An average value of elapsed
time is also presented on the lower axis of the altitude
scale. The lower segment of each chart provides the relationship between fuel remaining and distance traveled resulting from the flight profile shown. This portion may be
utilized to check actual inflight range data to provide assurance that adequate range is being achieved. The chart is
divided into 3 regions of Adequate Range, Inadequate
range-return to base, and Inadequate range-requiring emergency action. When an inflight range point is in the Adequate range region, the required mission range can be obtained by staying on the recommended flight profile.
However, the range may not be achieved if stronger headwinds are encountered as the flight progresses, and normal
pilot judgement must be used. These charts also assume
that the flight track is within proper navigational limits.
Standard temperature variation with PA is shown on the
upper segment of the charts. A general correction for temperature variation is to decrease IAS by 2.5 KTS and total
distance traveled by 0.5% for each 10°C above standard.
Detailed flight planning must always be made for the actual
aircraft configuration, fuel load, and flight conditions when
maximum range is required. This data is based on JP-4 fuel.
It can be used with JP-5, JP-8, aviation gasoline, or any
other fuels ONLY IF THE TAKEOFF GROSS WEIGHT
AND THE FUEL LOAD WEIGHT MATCH THE DATA
AT THE TOP OF THE CHART. The Flight Time and the
Distance Traveled data SHOULD NOT be used with any
full tank configuration if the fuel density is not approximately 6.5 lb/gal (JP-4 fuel).
a. SELF-DEPLOYMENT MISSION. The selfdeployment mission is shown in Figure 7A-40 and the
ESSS is configured with two 230-gallon tanks outboard and
two 450-gallon tanks inboard. In this configuration, the aircraft holds in excess of 11,000 lb of JP4 fuel and has a
take-off gross weight of 24,500 pounds in order to achieve
the desired mission range of 1,150 Nm. This gross weight
is allowed for ferry missions only, requiring low load fac-

7A-148

tors and less than 30 degree angle banked turns. This mission was calculated for a standard day with a constant 10
knot headwind added to be conservative. Since there may
not be any emergency landing areas available, the mission
should not be attempted if headwinds in excess of 10 knots
are forecast. Take-off must be made with a minimum of
fuel used (60 pounds) for engine start and warm-up, and a
climb to 2,000 feet should be made with maximum power
and airspeed between 80 and 105 KIAS. The first segment
should be maintained at 2,000 feet and 105 KIAS for 2
hours. The average engine TRQ should be about 79% for
this segment, but will initially be a little more and gradually
decrease. Altitude is increased in 2,000 feet increments to
maintain the optimum altitude for maximum range to account for fuel burn. The first 2 segments are for 2 hours
each, followed by 1 hour segments until reaching 10,000
feet. At this altitude, the airspeed for best range should also
be reduced to 100 KIAS for the remainder of the flight.
Engine bleed air was assumed to be off for this mission
except for that required for fuel tank pressurization. Electrical cabin heat may be used. Removal of the HIRSS
baffles (benign mode) will reduce fuel flow by about 14
lb/hr. If oxygen is available, continuation of the staircase
climb sequence to 15,500 feet PA will result in about 23
additional Nm of range capability.
b. ASSAULT MISSION PROFILE - 4 tanks. The assault mission profile is shown in Figure 7A-41 with the
ESSS configured with four 230-gallon tanks. In this configuration, the aircraft holds in excess of 8,300 pounds of
JP4 fuel and assumes a take-off gross weight of 22,000
pounds which provides a maximum mission range of 980
Nm with 400 lbs reserve. This mission was calculated for a
standard day with a zero headwind. Take-off must be made
with a minimum of fuel used (80 pounds) for engine start
and warm-up, and a climb to 4,000 feet should be made
with maximum power and airspeed between 80 and 108
KIAS. The first segment should be maintained at 4,000 feet
and 108 KIAS for 1 hour. The average engine TRQ should
be about 79% for this segment, but will initially be a little
more and gradually decrease. Altitude is increased in 2,000
feet increments to maintain the optimum altitude for maximum range to account for fuel burn. The segments are for 1
hour each, until reaching 10,000 feet. At this altitude, the
airspeed for best range should be reduced to 100 KIAS for
the remainder of the flight.

TM 1-1520-237-10

c. ASSAULT MISSION PROFILE – 2 tanks. The assault mission profile is shown in Figure 7A-42 with the
ESSS configured with two 230-gallon tanks. In this configuration, the aircraft holds in excess of 5,300 pounds of
JP4 fuel and assumes a take-off gross weight of 22,000
pounds which provides a maximum mission range of 580
Nm. with 400 lbs reserve. This mission was calculated for a
standard day with a zero headwind. Take-off must be made
with a minimum of fuel used (80 lbs) for engine start and
warm-up, and a climb to 4,000 feet should be made with

maximum power and airspeed between 80 and 108 KIAS.
The first segment should be maintained at 4,000 feet and
108 KIAS for 1 hour. The average engine TRQ should be
about 77% for this segment, but will initially be a little
more and gradually decrease as shown on each segment.
Altitude is increased in 2,000 feet increments to maintain
the optimum altitude for maximum range to account for
fuel burn. At this altitude, the airspeed for best range should
also be reduced to 100 KIAS for the remainder of the flight.

7A-149

TM 1-1520-237-10

EXAMPLE:
WANTED:
Assurance of adequate aircraft range for mission defined.
KNOWN:
Flight position: 300 nm from base
Flight Track Within Limits
Fuel Remaining = 7,900 pounds
Elapsed flight time = 2 HRS, 50 MINS (2.83 HRS)
Target: Normal Flight Conditions:
Airspeed = 105 KIAS
Press Alt = 2,000 feet
Approx Torque = 74%
METHOD:
(1)

(2)

To determine target nominial flight conditions, enter upper chart at elapsed flight time and move up
to determine target airspeed, approximate torque, and pressure altitude.

Figure 7A-40. Self Deployment Mission Profile (Sheet 1 of 2)

7A-150

TM 1-1520-237-10

SELF DEPLOYMENT MISSION PROFILE

ESSS/2−230 AND 2−450 GALLON TANK CONFIGURATION
STANDARD DAY 10 KT HEADWIND
HIRSS SUPPRESSED MODE
GROSS WEIGHT = 24,500 LB FUEL LOAD = 11,000 LB (JP4)
BLEED AIR OFF
60 LB WARM UP
−9

105 KIAS

100 KIAS
(~ 72%)

−5

(~ 56%)

−1
(~ 72%)
APPROX TRQ (~%)

3
(~ 72%)

7
(~ 74%)

STANDARD TEMP ~ oC

PRESSURE ALT ~ 1000 FT

(RECOMMENDED AIRSPEEDS)

11
(−79)

15
0

ELAPSED FLIGHT TIME ~ HRS

12000
(APPROX FULL FUEL)
11000

10000

9000

8000

(7900 LBS)

FUEL REM. ~ LBS

DESIRED MISSION RANGE

7000
ADEQUATE RANGE
6000

INADEQUATE RANGE
ABORT MISSION

5000

4000

3000

N
IO

INADEQUATE RANGE

2000

1000
LOW FUEL LIGHTS

0
0

DATA BASIS:

100

FLIGHT TEST

200

300
1

400

600

800

1000

DISTANCE TRAVELED ~ NM

1200

AA1030A
SA

Figure 7A-40. Self Deployment Mission Profile (Sheet 2 of 2)

7A-151

TM 1-1520-237-10

ASSAULT MISSION PROFILE FOR MAX RANGE
ESSS/4−230 GALLON TANK CONFIGURATION
STANDARD DAY ZERO HEADWIND
HIRSS SUPPRESSED MODE
GROSS WEIGHT = 22000 LB FUEL LOAD = 8280 LB (JP4)
80 LB WARM UP
BLEED AIR OFF
12

−9
(RECOMMENDED AIRSPEEDS ~ KIAS)

108

104

100
−5
(~ 56%)

(~ 71%)
8

−1
(~ 70%)

APPROX TRQ (~ %)

3
(~ 74%)

7
(~ 79%)

STANDARD TEMP ~ oC

PRESSURE ALT ~ 1000 FT

15
0

ELAPSED FLIGHT TIME ~ HRS
10000

9000
(APPROX FULL FUEL)
8000

FUEL REM. ~ LBS

7000
ADEQUATE RANGE
6000

5000

INADEQUATE RANGE
ABORT MISSION

4000

3000

N
IO

DI
RA

2000

INADEQUATE RANGE
1000
LOW FUEL LIGHTS

0
0

DATA BASIS:

200

FLIGHT TEST

400

600

800

1000

DISTANCE TRAVELED ~ NM

AA1031A
SA

Figure 7A-41. Assault Mission Profile (4 - 230 Gallon Tanks)

7A-152

TM 1-1520-237-10

ASSAULT MISSION PROFILE FOR MAX RANGE
ESSS/2−230 GALLON TANK CONFIGURATION
STANDARD DAY ZERO HEADWIND
HIRSS SUPPRESSED MODE
FUEL LOAD = 5280 LB (JP4)
GROSS WEIGHT = 22000 LB
80 LB WARM UP
BLEED AIR OFF
108

104

100

−5
(~ 70%)

(~ 65%)

−1

PRESS ALT

(~ 69%)
APPROX TRQ (~ %)
6

3
(~ 73%)

7
(~ 77%)

STANDARD TEMP ~ oC

(RECOMMENDED AIRSPEEDS ~ KIAS)

15
0

ELAPSED FLIGHT TIME ~ HRS

6000

5500
(APPROX FULL FUEL)
5000

4500

FUEL REM. ~ LBS

ADEQUATE RANGE

4000

3500

3000
INADEQUATE RANGE
ABORT MISSION
2500
S

2000

D
RA

1500

INADEQUATE RANGE
1000

500
LOW FUEL LIGHTS
0

DATA BASIS:

0
FLIGHT TEST

100

200

300

400

500

DISTANCE ~ NM

600

AA1032A
SA

Figure 7A-42. Assault Mission Profile (2 - 230 Gallon Tanks)

7A-153/(7A-154 Blank)

TM 1-1520-237-10

CHAPTER 8
NORMAL PROCEDURES
Section I MISSION PLANNING
8.1 MISSION PLANNING.

c. The crew chief will perform all duties as assigned by
the pilot.

Mission planning begins when the mission is assigned
and extends to the preflight check of the helicopter. It includes, but is not limited to checks of operating limits and
restrictions; weight, balance, and loading; performance;
publications; flight plan; and crew and passenger briefings.
The pilot in command shall ensure compliance with the
contents of this manual that are applicable to the mission
and all aviation support equipment required for the mission
(e.g., helmets, gloves, survival vests, survival kits, etc).
8.2 EH-60A DATA.

AN/ALQ-151(V)2; mission operator duties, equipment
checks, system initialization procedures, and self-test procedures are outlined in TM 32-5865-012-10.
8.3 AVIATION
(ALSE).

LIFE

SUPPORT

EQUIPMENT

8.5 CREW BRIEFING.
A crew briefing shall be conducted to ensure a thorough
understanding of individual and team responsibilities. The
briefing should include, but not be limited to, pilots, crew
chief, mission equipment operator, ground crew responsibilities, and the coordination necessary to complete the
mission in the most efficient manner. A review of visual
signals is desirable when ground guides do not have direct
voice communications link with the crew.
8.6 PASSENGER BRIEFING.
The following guide may be used in accomplishing required passenger briefings. Items that do not pertain to a
specific mission may be omitted.
a. Crew introduction.

All aviation life support equipment required for mission;
e.g., helmets, gloves, survival vests, survival kits, etc., shall
be checked.

b. Equipment.
(1) Personal, to include ID tags.

8.4 CREW DUTIES/RESPONSIBILITIES.
(2) Professional.
The minimum crew required to fly the helicopter is two
pilots. Additional crewmembers, as required, may be added
at the discretion of the commander. The manner in which
each crewmember performs his related duties is the responsibility of the pilot in command.

(3) Survival.
c. Flight data.
(1) Route.

a. The pilot in command is responsible for all aspects of
mission planning, preflight, and operation of the helicopter.
He will assign duties and functions to all other crewmembers as required. Prior to or during preflight, the pilot will
brief the crew on items pertinent to the mission; e.g., performance data, monitoring of instruments, communications,
emergency procedures, taxi, and load operations.

(2) Altitude.
(3) Time en route.
(4) Weather.
d. Normal procedures.

b. The pilot in command must be familiar with pilot
duties and the duties of the other crew positions.

(1) Entry and exit the helicopter.

8-1

TM 1-1520-237-10

(2) Seating.

(10) Weapons.

(3) Seat belts.

(11) Protective masks.

(4) Movement in helicopter.
(5) Internal communications.

(12) Parachutes.
(13) Hearing protection.
(14) Aviation life support equipment (ALSE).

(6) Security of equipment.
e. Emergency procedures.
(7) Smoking.
(1) Emergency exits.
(8) Oxygen.
(9) Refueling.

8-2

(2) Emergency equipment.
(3) Emergency landing/ditching procedures.

TM 1-1520-237-10

Section II OPERATING PROCEDURES AND MANEUVERS
8.7 OPERATING PROCEDURES AND MANEUVERS.
This section deals with normal procedures and includes
all steps necessary to ensure safe and efficient operation of
the helicopter from the time a preflight begins until the
flight is completed and the helicopter is parked and secured.
Unique feel, characteristics, and reaction of the helicopter
during various phases of operation and the techniques and
procedures used for taxiing, takeoff, climb, etc., are described, including precautions to be observed. Your flying
experience is recognized; therefore, basic flight principles
are avoided. Only the duties of the minimum crew necessary for the actual operation of the helicopter are included.
Additional crew duties are covered as necessary in Section
I Mission Planning. Mission equipment checks are contained in Chapter 4 Mission Equipment. Procedures specifically related to instrument flight that are different from normal procedures are covered in this section, following
normal procedures. Descriptions of functions, operations,
and effects of controls are covered in Section IV Flight
Characteristics, and are repeated in this section only when
required for emphasis. Checks that must be performed under adverse environmental conditions, such as desert and
cold-weather operations, supplement normal procedures
checks in this section and are covered in Section V Adverse
Environmental Conditions.
8.8 SYMBOLS DEFINITION.
Items which apply only to night or only to instrument
flying shall have an N or an I, respectively, immediately
preceding the check to which it is pertinent. The symbol O
shall be used to indicate 9if installed9. Those duties which
are the responsibility of the pilot not on the controls, will be
indicated by a circle around the step number; i.e., ④ . The
symbol star. indicates an operational check is required.
Operational checks are contained in the performance section of the condensed checklist. The asterisk symbol * indicates that performance of step is mandatory for all thruflights. The asterisk applies only to checks performed prior
to takeoff. Placarded items such as switch and control labels appear in uppercase type.
8.9 CHECKLIST.
Normal procedures are given primarily in checklist form,
and amplified as necessary in accompanying paragraph
form, when a detailed description of a procedure or maneu-

ver is required. A condensed version of the amplified
checklist, omitting all explanatory text, is contained in the
operator’s checklist. To provide for easier crossreferencing, the procedural steps in the checklist are numbered to coincide with the corresponding numbered steps in
this manual.
8.10 PREFLIGHT CHECK.
The pilot’s walkaround and interior checks are outlined
in the following procedures. The preflight check is not intended to be a detailed mechanical inspection. The preflight
order is a recommended sequence only. The expanded substeps do not need to be memorized or accomplished in order. The steps that are essential for safe helicopter operation are included. The preflight may be made as
comprehensive as conditions warrant at the discretion of
the pilot.
8.11 BEFORE EXTERIOR CHECK (FIGURE 8-1).

WARNING
Do not preflight until armament systems
are safe, switches off, safety pins installed
and locking levers in locked position.
1. Publications - Check; required forms and publications, and availability of operator’s manual(s) (-10) and checklist (-CL).
*2. Helicopter covers, locking devices, tiedowns,
and grounding cables - Removed and secured.
*3. Fuel - Check quantity as required.
4. Fuel sample - As required. Check for contamination before first flight of the day and after
adequate settling time after cold refueling, or if
fuel source is suspected contaminated.
8.12 EXTERIOR CHECK.
Exterior walkaround diagram is shown in Figure 8-1.

8-3

TM 1-1520-237-10

8.13 NOSE SECTION (AREA 1).

*3. Check main landing gear drag beam for cracks.
8.15 CABIN TOP (AREA 3).

CAUTION

1. Cabin top - Check as follows:
a. Left engine - Check inlet.

Do not deflect main rotor blade tips more
than 6 inches below normal droop position when attaching tiedowns. Do not tie
down below normal droop position.

b. Left pitot tube - Check.
c. Control access - Check flight controls, hydraulic reservoir, and filter indicators.
Check tempilabels for safe indication and
security. Check area.

*1. Main rotor blades - Check.
2. Fuselage - Nose area, check as follows:

d. Control access cover - Close and check secured.

a. Windshield and wipers - Check.
O

b. Blade deice OAT sensor, FAT indicator
probe(s) - Check.

e. Right pitot tube - Check.
f. Right engine - Check inlet.

c. Avionics compartment - Check equipment
as required; secure door.
O

g. IRCM - Check.

d. Antennas - Check.
2. APU - Check; oil level, use dipstick.
e. Landing and searchlights - Check.
O

3. APU IPS - Check.

8.14 COCKPIT - LEFT SIDE (AREA 2).
4. Gust lock - Check.
1. Cockpit area - Check as follows:
5. Main transmission - Check; oil level.
a. Cockpit door - Check.
b. Copilot seat, belts, and harness - Check.
c. FM and

antennas - Check.

d. Landing gear support fairing and step Check.

*6. Main rotor system - Check controls, dampers,
head, and blades. BIMt indicators - Check for
safe indication (yellow color).
8.16 INTERIOR CABIN (AREA 4).
1. Cabin - Check as follows:
a. Fire extinguishers - Check.

e. Position light - Check.
b. First aid kits - Check.
f. Main landing gear - Check.
O

HSS, VSP, ejector rack locking levers
locked, fairings, and external tanks Check; refueling caps secure.
ES

c. Pilot’s and copilot’s tilt-back release levers
- Lock position.
d. Cabin interior - Check security of stowed
equipment.

h. Gunner’s window - Check.
e. Cabin seats and belts - Check.
i. Ambient sense port - Check.
*2. Left engine oil level - Check.
8-4

Change 10

2. APU accumulator pressure gage - Check minimum 2,800 psi.

TM 1-1520-237-10

3. Transmission oil filter impending bypass indicator - Check.
4. Cargo hook:

b. Electrical connections condition and security.
5. Survival gear and mission equipment - Check
as required.

a. General condition and security.

Change 8

8-4.1/(8-4.2 Blank)

TM 1-1520-237-10

1
6

AREA 1
AREA 2
AREA 3
AREA 4

NOSE SECTION
COCKPIT − LEFT SIDE
CABIN TOP
INTERIOR CABIN

AREA 5
AREA 6
AREA 7
AREA 8

FUSELAGE − LEFT SIDE
TAIL PYLON
FUSELAGE − RIGHT SIDE
COCKPIT − RIGHT SIDE

AA0672A
SA

Figure 8-1. Exterior Check Diagram
8.17 FUSELAGE - LEFT SIDE (AREA 5).
1. Fuselage - Check as follows:
a. Cabin door - Check.
O*b.

VOL

Armament system - Check.

k. Tail landing gear - Check.
*2. Intermediate gear box - Check; oil level.
8.18 TAIL PYLON (AREA 6).
1. Tail pylon - Check as follows:
a. Tail pylon - Check.

c. Fuel tank filler ports - Check; caps secure,
doors secured.
d. External pneumatic inlet port - Door secured.
e. Engine exhaust - Check.
O

f. APU IPS exhaust - Check.
g. APU exhaust - Check.

h. Chaff, flare dispensers - Check; number
and programmer settings.
i. Lower anticollision light - Check.
j. Antennas - Check.

b. Stabilator - Check.
O

c. Radar detector and

antennas - Check.

d. Position light - Check.
e. Upper anticollision light - Check.
*2. Tail rotor - Check.
*3. Tail rotor gear box - Check; oil level.
8.19 FUSELAGE - RIGHT SIDE (AREA 7).
1. Fuselage - Check as follows:
a. Antennas - Check.

Change 10

8-5

TM 1-1520-237-10

EH
Aft avionics compartment circuit
breakers and ECS fluid level - Check.

8.21 BEFORE STARTING ENGINES.
NOTE

c. Fire bottles thermal plug - Check.
d. Engine exhaust - Check.

Before engine operation can be performed
with the gust lock engaged, all main rotor tie
downs shall be removed.

e. Fuel tank gravity filler port - Check cap
secure; door secured.

*1. Copilot’s collective - Extended and locked.

O*f.

VOL

Armament system - Check.

g. Cabin door - Check.
8.20 COCKPIT - RIGHT SIDE (AREA 8).
*1. Right engine oil level - Check.

2. Shoulder harness locks - Check.
3. PARKING BRAKE - Release, then set.
.4. Circuit breakers and switches - Set as follows:
a. Circuit breakers - In.

2. Cockpit area - Check as follows:

b. Avionics - Off, frequencies set.

c. BLADE DEICE POWER switch - OFF.

a. Ice detector - Check.
b. Ambient sense port - Check.

HSS, VSP, ejector rack locking levers
locked, fairings, and external tanks Check; refueling caps secure.

*d. Radar altimeter - Set.
feet.

Left LO bug 200

e. Clocks - Set and running.
f. BACKUP HYD PUMP - AUTO.

d. Gunner’s window - Check.
e. External electrical power receptacle - Door
secured.

*g. ANTICOLLISION/POSITION LIGHTS
- As required.
*h.

Q/F PWR switch - OFF.

O*i.

ECS panel switches - OFF.

f. Main landing gear - Check.
g. Position light - Check.
h. Landing gear support fairing and step Check.
i. FM and

antennas - Check.

j. Cockpit door - Check.

j. CARGO HOOK EMERG REL switch OPEN, ARMING switch - SAFE.
k. APU CONTR switch - OFF; APU
T-handle - In.
l. GENERATORS NO. 1 and NO. 2
switches - Check ON.

k. Pilot seat, belt, and harness - Check.
m. Ground power unit - Connected if required.
l. Set switch on dimmer control box as desired. NORM for IR dimming.

*n. AIR SOURCE HEAT/START switch APU (OFF for external air source).

*3. Check main landing gear drag beam for cracks.
o. EMER OFF T-handles - Full forward.
.*4. Crew and passenger briefing - Complete as required.
8-6

Change 10

*p. BATT switch - ON.

TM 1-1520-237-10

8.22 COCKPIT EQUIPMENT CHECKS.
*1. FUEL PUMP switch - APU BOOST.

caution/advisory panel BRT/DIM-TEST
switch is set to TEST. (This can also occur
in flight). The original indications may be
restored by pressing the applicable switches.

*2. APU CONTR switch - ON.
a. Caution/advisory panel BRT/DIM-TEST
switch - TEST. Caution/advisory/warning,
CIS/MODE SEL, and VSI advisory lights
on. #1 and #2 FUEL LOW caution lights
flashing. AFCS FAILURE ADVISORY
lights will illuminate. EH SYSTEMS SELECT switches will illuminate. EH ASE
advisory light - Press to test.

NOTE
If the APU does not start and the APU ACCUM LOW advisory light is not on with
the APU CONTR switch ON, the manual
override lever on the accumulator manifold
should be pulled to attempt another start, and
held until the APU has reached selfsustaining speed.
If APU fails, note and analyze BITE indications before cycling BATT switch or before
attempting another APU start.

WARNING

b. INSTR LT PILOT FLT control - ON.

c. Caution/advisory BRT/DIM-TEST switch
- BRT/DIM momentarily and then to
TEST.

d. All caution/advisory/warning panels CIS/
MODE SEL and VSI advisory lights on at
decreased intensity. AFCS FAILURE
ADVISORY lights will not dim.

Stabilator will move to full trailing edge
down position upon application of AC
power. Assure stabilator area is clear
prior to energizing stabilator system.

CAUTION

*3. APU generator switch - ON.
*4. EXT PWR switch - OFF and cable disconnected.
O.5.

FUEL
ERFS AUXILIARY
MENT control panel - TEST.

MANAGE-

O*6.

FUEL MANAGEERFS AUXILIARY
MENT control panel - Set fuel as required.

*7.

IINS SYSTEMS SELECT switches - DG
and VG.

If DEC signal validation codes are displayed on the % TRQ indicator, do not
fly the helicopter.
10.

701C DEC engine fault indicator codes - Check
for signal validation as required.

N 11. Interior/exterior lighting - Set.
O.12. Mission equipment - Check.

.*8.

.*13. Cold weather control exercise - Check if temperature is below -17°C (1°F).

IINS - Align.

9. Caution/advisory/warning panels - Check as required.

*14. AFCS FAILURE ADVISORY lights - If on,
POWER ON RESET.

NOTE

*15. SAS1 off, SAS2, TRIM, FPS, and BOOST
switches - Push ON.

Pulsating of any caution/advisory lights in
unison with the LOW ROTOR RPM warning lights may occur in the DIM mode.

.16. Flight controls - Check first aircraft flight of
day as follows:

The switch legend on the VSI/HSI and CIS
mode select panels may change when the

a. Collective - Midposition, pedals centered,
friction off.

Change 8

8-7

TM 1-1520-237-10

b. BOOST switch - Press off. There will be a
slight increase in collective and pedal
forces. BOOST SERVO OFF caution and
MASTER CAUTION lights should be on.

TAIL RTR SERVO ON advisory light illuminate. Move pedals through full range
in no less than 5 seconds. There should be
no binding.

c. Right SVO OFF switch - 1ST STG. No
allowable cyclic stick jump. #1 PRI
SERVO PRESS caution and MASTER
CAUTION lights should be on.

k. TAIL SERVO switch - NORMAL. Caution and advisory lights out.

d. Move cyclic and pedals slowly through full
range. There should be no binds or restrictions. Move collective full up to full down
in about 1 to 2 seconds. Check #2 PRI
SERVO PRESS caution light does not illuminate during movement of collective.

l. BOOST switch - ON. BOOST SERVO
OFF caution light should be off.
.17. Stabilator - Check.

WARNING

e. Right SVO OFF switch - 2ND STG. No
allowable cyclic stick jump. #2 PRI
SERVO PRESS caution and MASTER
CAUTION lights should be on.

If any part of stabilator check fails, do not
fly helicopter.

f. Repeat step d. above. Check #1 PRI
SERVO PRESS caution light does not illuminate during movement of collective.

For the purpose of this check, the right
STAB POS indicator shall be used. The left
STAB POS indicator may vary from right
indicator as much as 62° throughout the
check.

WARNING
If #1 PRI SERVO PRESS or #2 PRI
SERVO PRESS caution light illuminates
during collective movement, a servo bypass valve may be jammed. If this situation occurs, do not fly the helicopter.

NOTE

a. STAB POS indicator should be between
34° and 42° DN.
b. TEST button - Press and hold. Check
STAB POS indicator moves up 5° to 12°.
MASTER CAUTION and STABILATOR caution lights on; stabilator audio
heard.

g. SVO OFF switch - Center.
NOTE
During steps h. and i., check for not more
than 1.5 inches of freeplay in control.
h. Collective - Move through full range in no
less than 5 seconds. There should be no
binding.
i. Pedals - Move both pedals through the full
range in no less than 5 seconds. There
should be no binding.
j. TAIL SERVO switch - BACKUP. #1
TAIL RTR SERVO caution light, both
MASTER CAUTION lights, and #2
8-8

Change 8

c. AUTO CONTROL RESET switch Press ON. Note that the STABILATOR
caution light and audio are off, and STAB
POS indicator moves to 34° to 42° down.
d. Either cyclic mounted stabilator slew-up
switch - Press and hold until STAB POS
indicator moves approximately 15° trailing
edge up, release, stabilator should stop.
STABILATOR caution and MASTER
CAUTION lights on and beeping audible
warning in pilot’s and copilot’s headsets.
MASTER CAUTION - Press to reset audio tone.
e. Other cyclic mounted stabilator slew-up
switch - Press and hold until STAB POS

TM 1-1520-237-10

indicator moves approximately 15° trailing
edge up, release, stabilator should stop.

g. MAN SLEW switch - DN and hold until
STAB POS indicator reads 0°.

f. MAN SLEW switch - UP and hold until
stabilator stops. STAB POS indicator
should be 6° to 10° up.

Change 8

8-8.1/(8-8.2 Blank)

TM 1-1520-237-10

h. AUTO CONTROL RESET switch Press ON. STAB POS indicator should
move 34° to 42° DN. STABILATOR caution light off.
*18. Avionics - On.
NOTE
Only use map datums WGS-84 and NAD27. Other map datums were not verified using the Aviation Mission Planning System
(AMPS), and should not be used.
*18.1. Doppler/GPS - Program.
*18.2. Doppler/GPS mode select switch - OFF.
*19. COMPASS switch - SLAVED. Set as required.
20. Barometric altimeters - Set.
*21. Cyclic and pedals centered. Collective raise no
more than 1 inch (to prevent droop stop pounding) and friction.
22. BACKUP HYD PUMP switch - OFF.
O.23. Blade deice system - Test as required.

NOTE
PWR MAIN RTR and PWR TAIL RTR
monitor lights may flicker during tests in
steps e. through q.
c. BLADE DE-ICE TEST panel select
switch - NORM.
c.1. PWR MAIN RTR and TAIL RTR monitor lights - Press to test.
d. BLADE DEICE POWER switch - TEST.
e. PWR MAIN RTR and TAIL RTR monitor lights - Check. MAIN RTR monitor
light may go on for 2 to 4 seconds. If either
light remains on for 10 seconds or more:
(1) BLADE DEICE POWER switch OFF. If either light is still on:
(2) APU generator switch and/or EXT
PWR switch - OFF.
f. TEST IN PROGRESS light - Check. The
light should be on for 105 to 135 seconds.
No other blade deice system lights should
be on. PWR MAIN RTR and TAIL RTR
monitor lights may go on momentarily near
end of test. The TEST IN PROGRESS
light should then go off.

CAUTION

WARNING
Do not perform blade deice test when
blade erosion kit is installed.
To prevent overheating of droop stops,
blade deice test shall not be done more
than one time within a 30-minute period
when rotor head is not turning.
a. Ice rate meter PRESS TO TEST button Press and release.
b. Ice rate meter indicator - Moves to half
scale (1.0) holds about 50 seconds; then
falls to 0 or below. ICE DETECTED caution and MASTER CAUTION lights on
after 15 to 20 seconds into the test, and
FAIL flag should not be visible in flag window. Ice rate meter should move to zero
within 75 seconds after pressing PRESS
TO TEST button.

Droop stop hinge pins and cams may become very hot during test. Use care when
touching those components.
g. Crewman touch each droop stop cam Cams should be warm to touch.
h. BLADE DEICE POWER switch - OFF.
i. BLADE DE-ICE TEST panel select
switch - SYNC 1.
j. BLADE DEICE POWER switch - TEST.
MR DE-ICE FAIL caution and MASTER
CAUTION lights on.
k. BLADE DEICE POWER switch - OFF.
MR DE-ICE FAIL caution and MASTER
CAUTION lights off.

Change 10

8-9

TM 1-1520-237-10

l. BLADE DE-ICE TEST panel select
switch - SYNC 2.
m. BLADE DEICE POWER switch - TEST.
MR DE-ICE FAIL caution and MASTER
CAUTION lights on.
n. BLADE DEICE POWER switch - OFF.
MR DE-ICE FAIL caution and MASTER
CAUTION lights off.
o. BLADE DE-ICE TEST panel select
switch - OAT.
p. BLADE DEICE POWER switch - TEST.
MR DE-ICE FAIL caution, TR DE-ICE
FAIL caution, and MASTER CAUTION
lights on.
q. BLADE DEICE POWER switch - OFF.
MR DE-ICE FAIL caution, TR DE-ICE
FAIL caution, and MASTER CAUTION
lights off.
r. BLADE DE-ICE TEST panel select
switch - NORM.
24. Avionics - Check as required.
O . *25.

AFMS Auxiliary Fuel Management Panel TEST, set as required.

8.23 STARTING ENGINES.
*1. ENG FUEL SYS selector(s) - As required.
XFD for first start of day.
*2. Deleted.
*3. ENGINE IGNITION switch - ON.

ON. Complete EMER ENG SHUTDOWN
procedure.
a. If any of these indications occur, perform
EMER ENG SHUTDOWN as required.
(1) No TGT TEMP increase (light off)
within 45 seconds.
(2) No ENG OIL PRESS within 45 seconds.
(3) No % RPM 1 or 2 within 45 seconds.
(4) ENGINE STARTER caution light
goes off before reaching 52% Ng
SPEED.
(5) TGT TEMP reaches 700 850°C or
701C 851°C before idle is attained
(Ng 63%).

CAUTION

To avoid damage to the engine start
switch actuators, do not move the ENG
POWER CONT lever from IDLE to OFF
while pressing the starter button.
During engine start and runup ensure
that cyclic is kept in neutral, collective no
more than one inch above full down, and
pedals centered until % RPM R reaches
50% minimum to prevent damage to antiflap bracket bushings.
b. Starter button(s) - Press until Ng SPEED
increases; release.

*4. GUST LOCK caution light - Off.
NOTE
*5. Fire guard - Posted if available.
*6. Rotor blades - Check clear.
.* 7. Engine(s) - Start as follows:

CAUTION

If start is attempted with ENGINE IGNITION switch OFF, do not place switch
8-10

Change 10

If an ENGINE STARTER caution light
goes off when the starter button is released,
and the ENG POWER CONT lever is
OFF, the start attempt may be continued by
pressing and holding the starter button until
52% to 65% Ng SPEED is reached; then
release button.
c. TGT TEMP - Check below 700 150°C or
80°C before advancing ENG
701C
POWER CONT levers.

TM 1-1520-237-10

d. ENG POWER CONT lever(s) - IDLE.
Start clock.
e. System indications - Check.
f. ENGINE STARTER caution light(s).
Check, off at 52% to 65% Ng SPEED. If

ENGINE STARTER caution light remains on after 65% Ng:
(1) ENG POWER CONT lever - Pull
out.

Change 10

8-10.1/(8-10.2 Blank)

TM 1-1520-237-10

If caution light remains on:

b. HYD LEAK TEST switch - RESET. The
lights in a. should go off.

(2) APU - OFF or engine air source remove as required.
* 8. If single-engine start was made, repeat step 7
for other engine.
* 9. Systems - Check.
a. Ng SPEED - 63% or greater and within
3% of each other.
b. % RPM - Check that % RPM 1 or 2 is
not in the range of 20% to 40% and 60% to
90%. Advance ENG POWER CONT lever(s) as required.
c. XMSN PRESS - Check.
d. ENG OIL PRESS - Check.
e. #1 and #2 HYD PUMP caution lights Check off.
* 10. BACKUP HYD PUMP switch - AUTO.

NOTE
If the backup pump is still running following
the hydraulic leak test, cycle the BACKUP
HYD PUMP switch to OFF then back to
AUTO.
. 12. Tail rotor servo transfer - Check.
a. BACKUP HYD PUMP switch - AUTO
with backup pump not running.
NOTE
Failure of the BACK-UP PUMP ON advisory light or the #2 TAIL RTR SERVO
ON advisory light indicates a failure in the
leak detection/isolation system.
b. TAIL SERVO switch - BACKUP. #1
TAIL RTR SERVO caution light on and
#2 TAIL RTR SERVO ON and
BACK-UP PUMP ON advisory lights on
within 3 to 5 seconds.

. 11. Hydraulic leak test system - Check as follows:
NOTE
EH It is normal for the IINS CDU screen to
blank momentarily during the hydraulic leak
test system check.

c. TAIL SERVO switch - NORMAL. #1
TAIL RTR SERVO caution light and #2
TAIL RTR SERVO ON advisory light
off. BACK-UP PUMP ON advisory light
remains on for approximately 90 seconds.
O 13. Deleted.

When performing the HYD LEAK TEST,
all leak detection/isolation system components are checked electrically. Manually
holding the HYD LEAK TEST switch in
the test position does not allow the leak
detection/isolation system to be checked automatically. It manually holds the circuits
open. The switch must be placed in the
TEST position and released.
a. HYD LEAK TEST switch - TEST. #1
TAIL RTR SERVO, BOOST SERVO
OFF, SAS OFF, #1 and #2 RSVR LOW,
BACK-UP RSVR LOW, and MASTER
CAUTION lights and #2 TAIL RTR
SERVO ON and BACK-UP PUMP ON
advisory lights on. During this check, it is
normal for the collective and pedals to
move slightly.

*14. Deleted.
8.24 ENGINE RUNUP.
*1. Flight controls - Hold.

WARNING
Restrict the rate of ENG POWER CONT
lever’s movement, when the tailwheel
lockpin is not engaged. Rapid application
of ENG POWER CONT levers can result
in turning the helicopter, causing personnel injury or loss of life.
* 2. ENG POWER CONT lever(s) - FLY.

Change 10

8-11

TM 1-1520-237-10

*3. Droop stops - Check out 70% to 75% RPM R.
* 4. #1 and #2 GEN caution lights - Off.

CAUTION

a. BLADE DE-ICE TEST select switch EOT.

EH During operation of the air conditioner system, the right cabin door should
remain closed. If opening is required, the
right cabin door should not remain open
for more than 1 minute.

b. BLADE DEICE MODE select switch MANUAL M.

*5.
O

5 minutes before doing the deice EOT
check, to prevent blade overheating. Do
not do the deice EOT check if OAT is
above 38°C (100°F).

5.1

c. BLADE DEICE POWER switch - ON.
d. TR DE-ICE FAIL caution and MASTER
CAUTION lights on after 15 to 30 seconds, and MR DE-ICE FAIL caution light
on after 50 to 70 seconds.

ECS panel switches - As desired.

AUX CABIN HEATER switch - As desired.
NOTE

e. BLADE DEICE POWER switch - OFF.
TR DE-ICE FAIL caution, MR DE-ICE
FAIL caution, and MASTER CAUTION
lights off.

Cabin temperature must be below 29°C
(84°F) for heat to go on, and above 10°C
(50°F) for the heater to shut off.

f. BLADE DE-ICE TEST select switch NORM.

The auxiliary cabin heater circuit is reset
when bleed air transitions to and from the
APU. Crew reset of the auxiliary cabin
heater switch may be required after switching to or from the APU.
*5.2. Engine warmup - Check if temperature is below -17°C (1°F).

NOTE
If helicopter engine was started using external air source and/or external ac power, the
APU must be started to do APU generator
backup check.

a. At temperatures between -17°C (1°F)
and -43°C (-45°F), warm engines at
IDLE for 3 minutes.

g. GENERATORS NO. 1 or NO. 2 switch OFF. Applicable GEN and MASTER
CAUTION lights on.

b. At temperatures between -43°C (-45°F)
and -54°C (-65°F), warm engines at
IDLE for 5 minutes.

h. BLADE DEICE POWER switch - ON.
Wait 30 seconds, no deice lights on.

NOTE

i. GENERATORS switch(es) - ON. Applicable GEN caution light(s) off.

ECS heater will operate with either backup
pump or windshield anti-ice operating, but
not with both at the same time.

j. BLADE DEICE POWER switch - OFF.
k. BLADE DEICE MODE select switch AUTO.

O. 6. DEICE EOT - Check as required.

CAUTION

*7. % TRQ 1 and 2 - Matched within 5%.
* 8.

In ambient temperatures above 21°C
(70°F), operate rotor at 100% RPM R for
8-12

Change 10

Q/F PWR switch - As desired.

* 9. FUEL PUMP switch - OFF.

TM 1-1520-237-10

dry river beds) may be deferred (maximum of 5
flight hours) until a suitable location is reached.

* 10. APU CONTR switch - OFF.
* 11. AIR SOURCE HEAT/START switch - As required.
* 12. ENG FUEL SYS selectors - As required.

*19. FUEL BOOST PUMP CONTROL switches ON (for all fuel types). Indicator lights check On.
O . 20. ERFS AFMS External extended range fuel
transfer - Check.

* 13. SAS 1 - ON.
NOTE

8.25 BEFORE TAXI.
If a consecutive or random route was programmed, do not cycle through the test position. Cycling through test will delete any
preprogrammed routes.
*13.1. Doppler/GPS mode select switch - As desired.
*14. Collective friction - As required.
NOTE
A slight amount of collective friction (approximately 3 pounds) should be used to
prevent pilot induced collective oscillations.
N O*15. HUD - Adjust brightness, mode, barometric altitude, pitch, and roll as necessary.
O * 16.

EH IINS NAVRDY light flashing - CDU mode
select switch to NAV.

O * 17.

IINS SYSTEMS SELECT switches IINS.

WARNING
When on the ground, the ejector rack
lock lever should be turned inward to allow the pilot visual confirmation from the
cockpit. Prior to flight, the ejector rack
lock lever must be in the unlock (vertical)
position to allow emergency jettisoning of
the tanks in flight.
ES

O*1.
O*2.

Ejector rack lock levers unlocked.

VOL Volcano jettison safety pins - Remove
and red arming levers to arm.

WARNING
EH

WARNING
Engine anti-ice bleed and start valve malfunction can cause engine flameout.

Ensure the chaff arm switch is in the
SAFE position before the chaff pin is removed.
O*3. Chaff, flare electronic module(s) safety pin(s) Remove.
*4. Chocks - Removed.

18. Engine Health Indicator Test (HIT)/Anti-Icing
Check - Accomplish. Refer to ENGINE
HEALTH INDICATOR TEST/ANTI-ICE
CHECK IN HELICOPTER LOG BOOK. HIT/
ANTI-ICE checks while operating in adverse
conditions (e.g., dust, desert, coastal beach area,

*5. Doors - Secure.
*6. PARKING BRAKE - Release.
* 7. TAIL WHEEL switch - As required.

Change 10

8-13

TM 1-1520-237-10

8. Wheel brakes - Check as required.
8.26 GROUND TAXI.

* 1. ENG POWER CONT levers - FLY.
* 2. Systems - Check.
* 3. Avionics - As required.

CAUTION

* 4. Crew, passengers, and mission equipment Check.
When performing these maneuvers, cyclic
inputs should be minimized to prevent
droop-stop pounding.
Landing and searchlight have less than
one foot ground clearance when extended.
Use caution when taxiing over rough terrain when landing light and/or searchlight
are extended.
Increase collective and place cyclic forward of neutral to
start forward movement. Minimize forward cyclic movement to prevent droop stop pounding. Reduce collective to
minimum required to maintain forward movement. Soft or
rough terrain may require additional collective pitch. The
use of excessive collective pitch during taxi, especially at
light gross weights, can cause the tailwheel to bounce.
Regulate taxi speed with cyclic and collective and control
heading with pedals. Use brakes as required.

8.29 TAKEOFF.

WARNING
If the stabilator has not begun trailing
edge up movement by 30 to 50 KIAS,
abort the takeoff.
Refer to the height-velocity diagram, Figures 9-2 and
9-3, for avoid areas. Since suitable landing areas are often
not available, operating outside avoid areas during takeoff
and climb will provide the highest margin of safety.
8.30 AFTER TAKEOFF.

WARNING

8.27 HOVER CHECK.
1. Systems - Check caution/advisory panel, CDU
and PDU(s) for normal indication.
2. Flight instruments - Check as required.
3. Power - Check. The power check is done by
comparing the indicated torque required to
hover with the predicted values from performance charts.

8.28 BEFORE TAKEOFF.

ERFS
AFMS Fuel transfer sequence
must be carefully planned and executed in
order to maintain CG within limits.

O. 1.

O. 3.

ERFS
AFMS Extended range fuel system
transfer - As required.
EH

ASE - As required.

VOL

Mine launch, post mine launch - As re-

quired.

WARNING
Pitot heat and anti-ice shall be on during
operations in visible moisture with ambient temperature of 4°C (39°F) and below.
Failure to turn on pitot heat in icing conditions can cause the stabilator to program trailing edge down during flight. If
this occurs, manually slew the stabilator
to zero degrees.

8-14

Change 10

8.31 BEFORE LANDING.
1. TAIL WHEEL switch - As required.
2. PARKING BRAKE - As required.
3. Crew, passengers, and mission equipment Check.

ASE check - As required.

TM 1-1520-237-10

EH ECM
ANTENNA switch - RETRACT. Check ANTENNA RETRACTED advisory light - On. ECM operator report antenna deployed light - Off.
EH

IINS TACAN - OFF.

8.32 LANDING.

b. Slope landing. The tailwheel should be locked
and the parking brake should be set. For slope
landings and all ground operations, avoid using
combinations of excessive cyclic and low collective settings. Where minimum collective is
used, maintain cyclic near neutral position and
avoid abrupt cyclic inputs. During nose-down
slope landings, low-frequency oscillations may
be eliminated by moving cyclic toward neutral
and lowering collective.

CAUTION

8.33 AFTER LANDING CHECK.
During roll-on landing aerodynamic
braking with aft cyclic is permitted with
the tail wheel contacting the ground. Once
the main wheels touchdown, the cyclic
must be centered prior to reducing collective. Excessive aft cyclic may cause droop
stop pounding and contact between main
rotor blades and other portions of the aircraft. Aerodynamic braking is prohibited
once the main landing gear touches down.
Use brakes to stop the aircraft.
NOTE
Because of the flat profile of the main transmission, pitching the helicopter nose up as
in hover, may cause a transient drop in indicated main transmission oil pressure, depending on degree of nose-up attitude.
a. Roll-on landing. A roll-on landing may be used
when the helicopter will not sustain a hover, to
avoid hovering in snow or dust, if tail rotor
control is lost, or when operating with one
heavy external tank.

CAUTION

When landing the EH-60A in a nose
downslope configuration, exercise extreme
caution to prevent the main rotor blades
from contacting the aft DF antennas.
When the main wheels contact the
ground, center the cyclic prior to reducing collective. The cyclic should be centered before the collective is placed in full
down to prevent possible rotor/airframe
contact. If droop stop contact is felt prior
to the main wheels touching the ground,
abort landing attempt.
EH

1. TAIL WHEEL switch - As required.
2. Exterior lights - As required.
3. Avionics/mission equipment - As required.

8.34 PARKING AND SHUTDOWN.
1. TAIL WHEEL switch - As required.
2. PARKING BRAKE - Set.

2.1. FUEL BOOST PUMP CONTROL switches OFF.
3. Landing gear - Chocked.
O 4. AUXILIARY FUEL MANAGEMENT control panel
FUEL XFR MODE
ERFS
switch - OFF. AFMS XFER MODE switch
- OFF.
O5.

AUXILIARY
FUEL
ERFS
AFMS
MANAGEMENT panel PRESS switch(es) Off.

O6.

VOL Volcano red arming levers - SAFE and
jettison safety pins install.

O 7.

Ejector rack locking levers - Locked.

WARNING
Ensure the chaff arm switch is in the
SAFE position before the chaff pin is removed.
O 8. Chaff, flare electronic module(s) safety pin(s) Install.

Change 10

8-15

TM 1-1520-237-10

EH IINS SYSTEMS SELECT switches - DG/
VG.

O 10.

O 11.

11.1

IINS - OFF.
ECS panel switches - OFF.

AUX CABIN HEATER switch - OFF.

12. SAS 1 - Off.
12.1

DPLR GPS MODE SEL switch - Off.

13. DEICE, PITOT, ANTI-ICE, HEATER and
EH Q/F PWR switches - OFF.
14. AIR SOURCE HEAT/START switch - APU.

20. ENGINE IGNITION switch - OFF.

21. Cyclic - As required to prevent anti-flap pounding.
22. Droop stops - Verify in, about 50% RPM R. If
one or more droop stops do not go in during
rotor shutdown, shut down an engine to lower
rotor idling RPM in an attempt to seat the droop
stops. If droops still do not go in, accelerate
rotor to above 75% RPM R. Repeat rotor shutdown procedures, slightly displacing cyclic in
an attempt to dislodge jammed droop stop. If
droop stops still do not go in, make certain that
rotor disc area is clear of personnel and proceed
with normal shutdown procedures while keeping cyclic in neutral position.

15. FUEL PUMP switch - APU BOOST.
CAUTION

16. APU CONTR switch - ON. The APU ON,
BACK-UP PUMP ON, and APU ACCUM
LOW advisory lights - On.

NOTE
If external electrical power is required for
shutdown, it shall be connected and EXT
PWR switch placed to RESET; then ON. If
external ac power is not available, complete
normal shutdown on right engine before
continuing.

To prevent damage to anti-flap stops, do
not increase collective pitch at any time
during rotor coast-down.
23. BACKUP HYD PUMP switch - OFF.

24. Stabilator - Slew to 0° after last flight of the
day.
25. BACK-UP PUMP ON advisory light - Check
off.

17. Collective raise no more than 1 inch.
18. Flight controls - Hold.

CAUTION

During shutdown ensure that cyclic is
kept in neutral or displaced slightly into
prevailing wind, collective no more than
one inch above full down and pedals centered.
Restrict the rate of ENG POWER CONT
lever movement, when the tailwheel lockpin is not engaged. Abrupt application of
ENG POWER CONT lever can result in
turning the helicopter.
19. ENG POWER CONT levers - IDLE.

8-16

Change 10

CAUTION

Before moving ENG POWER CONT lever OFF, engine must be cooled for 2
minutes at an Ng SPEED of 90% or less.
If an engine is shut down from a high
power setting (above 90%) without being
cooled for 2 minutes, and it is necessary to
restart the engine, the restart should be
done within 5 minutes after shutdown. If
the restart can not be done within 5 minutes, the engine should be allowed to cool
for 4 hours before attempting an engine
restart.
26. ENG POWER CONT levers - OFF after 2
minutes at Ng SPEED of 90% or less.
27. ENG FUEL SYS selectors - OFF.

TM 1-1520-237-10

28. TGT TEMP - Monitor. If TGT TEMP rises
above 538°C:

c. WINDSHIELD WIPER.
d. VENT BLOWER.

a. Start button - Press.
b. ENG POWER CONT lever(s) - Pull after
TGT TEMP is below 538°C.
O 29. Deleted.
30.

701C

DEC torque indicator fault code - Check.

31. Avionics - Off.
32. Deleted.
O33. HUD ADJ/ON/OFF switch - OFF.
34. Overhead switches - As required:
a. ANTICOLLISION/POSITION
LIGHTS.
b. Left panel light controls.

e. Right panel light controls.
35. APU generator switch - OFF.
36. FUEL PUMP switch - OFF.
37. APU CONTR switch - OFF.
38. BATT switch - OFF.
8.35 BEFORE LEAVING HELICOPTER.
1. Walkaround - Complete, checking for damage,
fluid leaks and levels.
2. Mission equipment - Secure.
3. Complete log book forms.
4. Secure helicopter - As required.

Change 10

8-16.1/(8-16.2 Blank)

TM 1-1520-237-10

Section III INSTRUMENT FLIGHT
8.36 INSTRUMENT FLIGHT.
Refer to FM 1-240 for instrument flying and navigation
techniques.

8-17

TM 1-1520-237-10

Section IV FLIGHT CHARACTERISTICS
8.37 GENERAL.

WARNING
Pedal trim switches must be pressed while
changing the helicopter heading during
hover. Do not hold hover heading against
yaw trim force. A rapid release of pedal
trim force will allow the FPS heading hold
feature to immediately correct to the last
known engaged heading. This can result
in rapid, divergent helicopter heading deviations.
a. Refer to FM 1-203 Fundamentals of Flight for explanation of aerodynamic flight characteristics.
b. The safe maximum operating airspeed range is described in Chapter 5. While hovering in high wind,
sideward and rearward flight should be limited to low
ground speeds. The helicopter is directionally stable in forward flight. In sideward and rearward flight, directional
control is more difficult. During approach, or slow flight as
the airspeed reaches about 17 to 20 KIAS, a mild vibration
will be felt.
8.38 GROUND RESONANCE.
Ground resonance is a self-excited vibration created
when a coupling interaction occurs between the movement
of the main rotor blades and the helicopter. For this to
happen, there must be some abnormal lead/lag blade condition which would dynamically unbalance the rotor and a
reaction between the helicopter and ground, which could
aggravate and further unbalance the rotor. Ground resonance can be caused by a blade being badly out of track, a
malfunctioning damper, or a peculiar set of landing conditions. Ground resonance may occur when a wheel reaction
aggravates an out-of-phase main rotor blade condition such
as a hard one-wheel landing, resulting in maximum lead
and lag blade displacement. This helicopter does not have a
history of ground resonance. If it should occur, get the helicopter airborne. If this is not possible, immediately reduce
collective pitch, place ENG POWER CONT levers OFF,
and apply wheel brakes.
8.39 MANEUVERING FLIGHT.
8.39.1 Flight with External Loads. Refer to FM 55450-1.
8-18

Change 10

WARNING
Static electricity generated by the helicopter should be discharged before attempting a sling or rescue hoist pickup. Use a
conductor between helicopter and the
ground to discharge the static electricity.
Caution must be exercised when transporting external loads that exhibit unstable characteristics. These loads may
amplify any oscillation and cause the load
to contact the aircraft.
a. Load bobble. In forward flight at the higher external
cargo hook load weights, a slight vertical bobble may occasionally be noticed. If experienced, this bobble will increase in amplitude with a corresponding increase in airspeed or aggressiveness of maneuver. This bobble is caused
by an external disturbance (e.g. turbulence or a control input) that triggers the natural elastic response of the sling.
To correct, airspeed shall be decreased or limit aggressiveness of maneuver until bobble is eliminated and pilot is
comfortable with the aircraft’s control.
b. Stabilator angle in level flight. Due to the increased
drag of external loads, collective position for a given level
flight speed will be higher. Correspondingly, the stabilator
angle will be more trailing edge down than usual. Since the
surface area and inherent drag of each external load varies,
exact guidance relative to how much more trailing edge
down angle that results is not possible.
c. Collective friction. With external cargo hook sling
loads, it is especially important to have collective friction
set at a minimum of three pounds.
8.39.2 Flying Qualities
Installed. ERFS

with

External

ERFS

a. Pitch Attitude vs. Airspeed. The ERFS installation
naturally results in increased drag. Since this drag vector is
below the center of gravity of the helicopter, the pitch attitude will be more nose-down for any speed beyond 60 to
70 KIAS. At mid to high gross weights (and most especially at a forward CG) there is a slight pitch down at 50 to
55 KIAS. The installation of the ERFS results in a small
increase in this nose-down tendency.

TM 1-1520-237-10

b. Tank Vibration. It will be observed that the right hand
tank(s) will vibrate more than the left tank(s). This is a
normal occurrence.

enced. During flight, if vertical oscillation is encountered,
the pilot should remove the hand from the collective grip;
this should eliminate the oscillation.

c. Stabilator Angle vs. Airspeed. With the increased
drag of the ERFS, a given airspeed will require more collective which, due to the collective to stabilator coupling,
results in a more trailing edge down stabilator angle. In the
ferry configuration (full inboard 450-gallon tanks, full outboard 230-gallon tanks) the stabilator angle at higher speeds
may be increased because of higher collective positions signal. This is normal as no stabilator program changes were
made for the ERFS.

8.39A 700 TRANSIENT ROTOR DROOP CHARACTERISTICS

d. Roll Attitude Hold (FPS ON). With only the ERFS
wings installed, the roll attitude hold feature of the FPS is
not noticeably affected. With full outboard 230-gallon tanks
there is a very slight degradation of roll attitude stability,
evidenced by a slower return to trim after an excitation
(gust). With four full 230-gallon tanks the return to trim is
a bit slower and with full inboard 450-gallon tanks and full
outboard 230-gallon tanks the return to trim is slower. Since
the return to trim is affected by the roll inertia of the helicopter, it is therefore recommended that for a four tank
configuration the outboard tanks be used first.

b. During descent with little or no collective applied,
Ng SPEED will be less than 80%. If % RPM R increases
above 100%, the ECU torque motor input to the HMU is
trimmed down in an attempt to restore 100% RPM 1/2 and
% RPM R. When collective is increased, the LDS input
demands more power, but the ECU continues to trim down
until % RPM 1/2 returns to 100%. Since the Ng SPEED is
at a slow speed, engine response time is greater. If rotor
drag increases faster than the engine controller response,
rotor droop occurs.

8.39.3 Collective Bounce/Pilot Induced Oscillation.
NOTE
The friction force refers to the breakaway
force required to move the collective stick in
an upward direction. The three pounds force
is measured with the BOOST servo and SAS
amplifiers operating and collective at midrange.
To prevent vertical oscillation (collective bounce), the
collective control system requires a minimum friction of
three pounds measured at the collective head. Vertical oscillation can occur in any flight regime and may be caused
by such events as SAS oscillation, turbulence, external load
oscillation, and inadvertent pilot input into the collective.
The oscillation causes the aircraft to vibrate. This vibration
will be felt as a vertical bounce at approximately three
cycles per second. If the severity of the oscillation is allowed to build, very high vibration levels will be experi-

a. The T700 engine control system accurately maintains
100 % RPM R throughout the flight envelope for most
maneuvers. However, pilots should be aware that certain
maneuvers performed with minimum collective applied will
result in significant transient rotor droop. High density altitudes, heavy gross weights and operation at less than
100% RPM R will aggrevate this condition.

c. During aggressive level deceleration (quick stop) or
right turn approach maneuvers as the collective is raised
and the nose lowered, % RPM R may droop to 95% or
lower for 1 - 2 seconds. % RPM R may then momentarily
increase to 105-106% as the engine control system overcompensates for the reduced % RPM 1/2. Similar conditions of low collective, high % RPM R, and low Ng
SPEED may be present during practice autorotations to a
power recovery. After the flare as the nose is leveled and
collective is increased, significant transient droop can occur. A rapid collective pull will aggravate the rotor droop.
d. Maneuvers that rapidly load the rotor system with no
collective input can result in transient droops as low as
92%. Transient droop is more pronounced at higher altitudes since the HMU reduces Ng SPEED acceleration as
barometric pressure decreases.
e. To minimize transient rotor droop, avoid situations
which result in rapid rotor loading from low Ng SPEED
and % TRQ conditions. Initiate maneuvers with collective
inputs leading or simultaneous to cyclic inputs. During approach and landing, maintain at least 15% - 20% TRQ and
transient droop will be minimal as hover power is applied.

Change 10

8-19

TM 1-1520-237-10

Section V ADVERSE ENVIRONMENTAL CONDITIONS
8.40.1.3 Inboard External Fuel Tanks or Stores.

8.40 GENERAL.
This section informs the crewmembers of the special
precautions and procedures to be followed during the various weather and climatic conditions that may be encountered. This will be additional material to that already covered in other chapters regarding the operation of the various
helicopter systems. Refer to FM 1-202 for cold weather
operations.
8.40.1 Shipboard Operations.
8.40.1.1 Helicopter Shutdown.

CAUTION

To avoid damage caused by slowing rotor
blades, crews shall exercise extreme caution when shutting down the helicopter on
board ships.
During rotor coast down on board ship, changing wind
conditions, gusts, flight deck turbulence, and rotor downwash from other helicopters can create excessive blade
flapping and cause helicopter damage.
8.40.1.2 Blade Folding.

CAUTION

Crews shall exercise extreme caution
when folding or spreading blades in high
wind/severe deck motion conditions.
Takeoff and landing adjacent to helicopters with folded blades should not be conducted on board ship. The folded blades
could contact each other and cause damage.
Folding or spreading of main rotor blades may be required on board ship to conserve flight deck space, to use
the ship’s elevator, or to hangar the helicopter. Controlling
the main rotor blades is significantly more difficult than on
land due to the effects of wind speed and direction, and ship
motion. High winds and/or ship motion can also cause
damage to folded blades.

8-20

Change 10

WARNING
During significant deck motion conditions, chocking must be accomplished expeditiously. On board ship operations will
expose flight deck personnel to risk of injury in the event of inadvertent jettison or
helicopter movement while chocking.
Inboard mounted external fuel tanks or stores significantly impede access to main wheels. Under severe deck
motion, the risk of helicopter movement may exceed the
risk of inadvertent stores jettison circuits prior to chocking.
Consideration should be given to not carrying inboard
mounted tanks or stores when severe deck motion is expected.
8.41 COLD WEATHER OPERATION.
The basic helicopter with normal servicing can operate
at temperatures down to −34°C (−29°F).

WARNING
Static electricity generated by the helicopter should be discharged before attempting a sling or rescue hoist pickup. In cold,
dry climatic conditions static electricity
buildups are large. Use a conductor between the helicopter and the ground to
discharge the static charge. Delay lowering rescue hoist hook until helicopter is
over the load, to lessen static charge
buildup.
NOTE
During operation in cold weather, particularly when snow or moisture is present, the
tail wheel locking indicating systems may
give erroneous cockpit indications.

TM 1-1520-237-10

8.41.1 Cold Weather Preflight Check.

times during 1 minute of control cycling in
step a.

CAUTION

Ice removal shall never be done by scraping or chipping. Remove ice by applying
heat or deicing fluid.
Blade deice operation with erosion strips
installed may cause blade damage.

(2) Move each tail rotor pedal alternately
through 3/4-inch of travel from neutral position 30 times during 1 minute of control
cycling in step a.
b. At temperatures between -31°C (-24°F) and
-43°C (-45°F), cycle collective slowly for 2
minutes.

a. In addition to the checks in Section II check aircraft
for ice or snow. If ice or snow is found, remove as much as
possible by hand and thaw aircraft with heated air or deicing fluid before attempting start. Failure to remove ice and
snow may cause damage.

(1) Move collective stick grip up about 1-1/2
inches from lower stop and down again
during first minute, and 3 inches of travel
during second minute of control cycling in
step b.

b. Check main rotor head and blades, tail rotor, flight
controls and engine inlets and hand holds for ice and snow.
Failure to remove snow and ice accumulations can result in
serious aerodynamic, structural effects in flight and serious
foreign object damage if ice is ingested into the engine.
Check ENG POWER CONT levers for freedom of movement.

(2) Move each tail rotor pedal alternately
through 3/8-inch of travel from neutral position during first minute and 3/4-inch of
travel during second minute of control cycling in step b.

c. On aircraft equipped with Extended Range Fuel System, check ESSS and 230/450-gallon fuel tank for ice or
snow. Remove as much as possible by hand and then use
heated air. Start APU and turn on pressure to both INBD
and OUTBD fuel tanks. Wing-mounted pressure regulator
may require heated air applied directly onto the exhaust
vent protruding from the ESSS wing. After regulator valve
is operating and fuel tanks are pressurized, leave system on.
DO NOT TURN OFF PRESSURE SWITCHES OR PRESSURE REGULATORS MAY FREEZE.

c. At temperatures between -43°C (-45°F) and
-54°C (-65°F), cycle collective stick grip slowly
for 5 minutes.
Move collective and pedals through travel for times
shown below:

Collective
Travel
(Approximately)

Pedals Travel
(Approximately)

Time
Duration

d. When parking the helicopter in temperatures below
freezing, the gust lock may seize due to frozen moisture in
rod assembly. Normal operations may be returned by
warming the assembly. Main rotor tiedowns may be used in
lieu of gust lock to meet parking requirements.

3/4-inch

1/8-inch

First minute

1-1/2 inches

1/4-inch

Second minute

1-3/4 inches

1/2-inch

Third minute

2-1/2 inches

5/8-inch

Fourth minute

8.41.2 Cold Weather Control Exercise. After starting the APU, the controls must be exercised when operating in a temperature range of -17°C (1°F) and below. The
control exercise is required

3 inches

3/4-inch

Fifth minute

a. At temperatures between -17°C (1°F) and
-31°C (-24°F), cycle collective control slowly
for 1 minute.
(1) Move collective stick grip up about 3
inches from lower stop, and down again 30

8.41.3 Engine Operation.
a. Even though cold weather does not particularly affect
the engine itself, it still causes the usual problems of ice in
the fuel lines, control valves, and fuel sumps, which frequently prevent a successful cold weather start. It may be
found that certain elements or accessories need preheating.

Change 10

8-21

TM 1-1520-237-10

warm-up will depend on temperature of the engine and lubrication system before start.
CAUTION

When starting an engine that has been exposed to low temperatures, watch for rise
in TGT TEMP within 45 seconds. If no
TGT TEMP rise is evident, manually
prime the engine and attempt another engine start. If there is no overboard fuel
flow during prime, inspect for ice in the
sumps and filters. During cold weather
operation, allow longer warm-up period
to bring transmission oil temperature up
to desired operating range refer to Chapter 5. Monitor oil pressure and temperature closely. When advancing the power
control levers, maintain transmission oil
pressure in normal operating range.
b. When starting in cold weather below -40°C (-40°F),
if light-off does not occur within 45 seconds after initial
indication of Ng SPEED, move ENG POWER CONT
lever for the affected engine back to OFF. With the engine
shutdown move the ENG POWER CONT lever from OFF
to FLY. If the force required to move the ENG POWER
CONT lever is higher than normal, suspect possible frozen
PAS cable. This situation may require maintenance prior to
attempting another start. If force is normal then attempt
another start. If light-off still does not occur within 45 seconds, abort start and do the following:
(1) ENG POWER CONT lever(s) - Hold at
LOCKOUT.
(2) FUEL BOOST PUMP CONTROL switch(es)
- ON until crewmember reports fuel from the
overflow drain.
(3) FUEL BOOST PUMP CONTROL switch(es)
- OFF.
(4) ENG POWER CONT lever(s) - OFF.
(5) Attempt another start.

b. During starts in extreme cold weather (near -54°C
(-65°F)) , the following oil pressure characteristics are typical:
(1) Oil pressure may remain at zero for the first 20
to 30 seconds after initiating the start. Abort the
start if oil pressure does not register within 1
minute after initiating a start.
(2) Once oil pressure begins to indicate on the
gage, it will increase rapidly and it will exceed
the limit. This condition is normal. The time for
oil pressure to decrease will depend on the ambient temperature, but should be normal within
5 minutes after starting the engine.
(3) Oil pressure may increase above the maximum
pressure limit if the engine is accelerated above
idle while oil temperature is within normal operating range. The pressure will decrease to
within the normal operating range as the oil
temperature increases.
c. It is normal for the OIL FLTR BYPASS caution
light to be on when starting an engine with oil temperatures
below normal because of high oil viscosity and the accumulation of oil filter contaminants. When the engine oil
temperature reaches about 38°C (100°F) during warm-up,
the light should go off.
8.41.5 Taxiing. The helicopter should not be taxied until
all engine temperatures and system pressures are within the
normal range. All taxiing should be done at low speeds
with wide-radius turns. If the tires are frozen to the surface,
a slight yawing motion induced by light pedal application
should break them free. Taxiing in soft snow requires
higher than normal power.
8.42 DESERT AND HOT WEATHER OPERATION.
Prolonged hovering flight in hot weather 35°C (95°F) at
higher gross weight may cause transmission oil temperature
to rise into the yellow precautionary range. Hovering operations in the precautionary range under those conditions
may be considered normal.

8.41.4 Engine Oil System Characteristics.
a. It is normal to observe high engine oil pressure during initial starts when the ambient temperature is 0°C
(32°F) or below. Run engine at idle until oil pressure is
within limits. Oil pressure should return to the normal range
after operating 5 minutes. However, time required for
8-22

Change 10

8.42.1 Taxiing and Ground Operation. Braking and
ground operation should be minimized to prevent system
overheating. During ground operations, if engine oil pressure falls into the red gage range when the power control
lever is in the idle position and/or the engine oil pressure
caution light comes on when the power control lever is in

TM 1-1520-237-10

the idle position, slightly advance the power control lever.
If the engine oil pressure returns to the yellow range and
the engine oil pressure caution light extinguishes, engine
oil pressure is acceptable.
8.43 IN-FLIGHT.
8.43.1 Thunderstorm Operation.

CAUTION

Avoid flight in or near thunderstorms, especially in areas of observed or anticipated lightning discharges.
a. Tests have shown that lightning strikes may result in
loss of automatic flight controls (including stabilator), engine controls or electrical power. The high currents passing
through the aircraft structure are expected to produce secondary effects whereby damaging voltage surges are
coupled into aircraft wiring.
b. If a lightning strike occurs whereby all aircraft electrical power and electronics subsystems and controls are

lost (including the engine 700 ECU/ 701C DEC and the
engine-driven alternator), both engines go immediately to
maximum power with no temperature limiter or overspeed
protection. In addition, the 701C engine overspeed may result in single or dual-engine shutdown without automatic
relight.
8.43.2 Turbulence.
a. Recommended maximum turbulence penetration airspeeds. For moderate turbulence, limit airspeed to the MAX
RANGE ( 700 Chapter 7 or 701C Chapter 7A) or Vne minus 15 knots, whichever is less.
b. In turbulent air - Maintain constant collective and use
the vertical situation indicator as the primary pitch instrument. The altimeter and vertical velocity indicator may vary
excessively in turbulence and should not be relied upon.
Airspeed indication may vary as much as 40 KIAS. By
maintaining a constant power setting and a level-flight attitude on the vertical situation indicator, airspeed will remain relatively constant even when erroneous readings are
presented by the airspeed indicator.

Change 10

8-22.1/(8-22.2 Blank)

TM 1-1520-237-10

8.43.3 Ice and Rain Operation.

CAUTION

Operation in rain will result in significant
damage to the blade erosion kit materials
and should be avoided.
At airspeeds greater than 120 KIAS or
during periods of reduced rain intensity
the windshield wipers may slow noticeably. If this occurs, wipers must be
parked immediately to avoid wiper motor
failure.
8.43.4 In-Flight Icing.

enced during normal operation of the blade deice system
because of ice build-up. The crew should closely monitor
engine instruments to prevent exceeding limits and/or rotor
droop. Significant power losses and increased fuel consumption will occur with the activation of engine inlet antiicing systems. Refer to Chapter 7 for torque available. The
main rotor hub and the blades collect ice before initiation of
a deice cycle. When enough ice has collected on the blades,
moderate vibration levels of short duration can be expected
in controls and airframe during normal deicing cycles. If
the blade deice system is not operating, unbalanced loads of
ice, resulting from asymmetric shedding, may cause severe
vibrations. However, these vibrations normally subside after 30 to 60 seconds when ice from other blades is shed.
d. ERFS AFMS When helicopter is equipped with
external extended range fuel system turn on pressure to
both INBD and OUTBD fuel tanks. This will prevent ice
accumulation and assure pneumatic pressure for fuel transfer.

CAUTION

NOTE
Activation of anti-ice systems after entry
into potential icing conditions creates the
possibility of engine FOD caused by ice
shedding. The ice detector has been designed primarily as a sensor to indicate
the requirement for activation of the
blade deice system.
a. All anti-ice systems must be turned on prior to entering visible moisture at ambient temperatures of 4°C (39°F)
or less.
b. If icing conditions are encountered, turn on all antiicing equipment immediately. If torque required increases
20% above that required for level flight at the airspeed
being maintained before entering icing, exit the icing environment or land as soon as possible. A 20% torque increase
indicates that normal autorotational rotor rpm may not be
possible, should dual-engine failure occur.
c. When the helicopter is equipped with an operating
blade deice, and icing conditions are encountered, a recurring torque increase up to 14% per engine may be experi-

After pressurizing the external extended
range fuel tanks, DO NOT TURN OFF if
ambient temperature is below 4°C (39°F).
8.43.5 Ground Operations.
a. Strong gusty winds may cause increased flapping of
the main rotor blades during shutdown following an icing
encounter, because the anti-flap restrainers may be frozen
in the fly position.
b. During flight in icing conditions when droop stop
heaters are not installed or fail to operate properly, the
droop stop hinges may become iced, resulting in the droop
stops not returning to the static position during rotor coast
down. When the droop stops do not return to the static
position, the main rotor blades may droop to within 4 feet
of the ground during shutdown. Strong gusty winds may
also cause excessive flapping of the main rotor blades, presenting the additional hazard of potential contact with the
aft fuselage. If the droop stops are suspected to be stuck in
the fly position, caution must be taken during shutdown to
be sure personnel remain clear of the helicopter.

Change 8

8-23/(8-24 Blank)

TM 1-1520-237-10

CHAPTER 9
EMERGENCY PROCEDURES
Section I AIRCRAFT SYSTEMS
9.1 HELICOPTER SYSTEMS.
This section describes the helicopter systems emergencies that may reasonably be expected to occur and presents
the procedures to be followed. Emergency operation of
mission equipment is contained in this chapter, insofar as
its use affects safety of flight. Emergency procedures are
given in checklist form when applicable. A condensed version of these procedures is contained in the condensed
checklist TM 1-1520-237-CL.

c. The term AUTOROTATE is defined as adjusting the
flight controls as necessary to establish an autorotational
descent and landing.
d. The term EMER ENG SHUTDOWN is defined as
engine shutdown without delay. Engine shutdown in flight
is usually not an immediate-action item unless a fire exists.
Before attempting an engine shutdown, identify the affected
engine by checking ENG OUT warning lights, % RPM,
% TRQ, ENG OIL PRESS, TGT TEMP, and Ng
SPEED.

9.2 IMMEDIATE ACTION EMERGENCY STEPS.
NOTE
The urgency of certain emergencies requires
immediate and instinctive action by the pilot. The most important single consideration
is helicopter control. All procedures are subordinate to this requirement. The MASTER
CAUTION should be reset after each malfunction to allow systems to respond to subsequent malfunctions. If time permits during
a critical emergency, transmit MAYDAY
call, set transponder to emergency, jettison
external stores if required, turn off boost
pumps, and lock shoulder harnesses.
Those steps that shall be performed immediately in an
emergency situation are underlined. These steps must be
performed without reference to the checklist. Nonunderlined steps should be accomplished with use of the checklist.
9.3 DEFINITION OF EMERGENCY TERMS.
For the purpose of standardization, these definitions shall
apply.
a. The term LAND AS SOON AS POSSIBLE is defined as landing at the nearest suitable landing area (e.g.,
open field) without delay. (The primary consideration is to
ensure the survival of occupants.)
b. The term LAND AS SOON AS PRACTICABLE is
defined as landing at a suitable landing area. (The primary
consideration is the urgency of the emergency.)

1. ENG POWER CONT lever(s) - OFF.
2. ENG FUEL SYS selector(s) - OFF.
3. FUEL BOOST PUMP CONTROL switch(es)
- OFF.

CAUTION

If TGT rises above 538°C after shutdown,
place AIR SOURCE HEAT/START
switch as required, turn ENGINE IGNITION switch OFF, and press starter to
motor engine for 30 seconds or until TGT
TEMP decreases below 538°C.
e. The term LOCKOUT is defined as manual control of
engine RPM while bypassing 700 ECU, or 701C DEC functions. Bypass of the engine control will be required when
% RPM 1 or 2 decreases below normal demand speed.

CAUTION

When engine is controlled with ENG
POWER CONT lever in LOCKOUT, engine response is much faster and TGT
limiting system is inoperative. Care must
be taken not to exceed TGT limits and
keeping % RPM R and % RPM 1 and 2
in operating range.

Change 8

9-1

TM 1-1520-237-10

ENG POWER CONT lever - Pull down and advance
full forward while maintaining downward pressure, then
adjust to set % RPM R as required. Engine control malfunctions can result in % RPM R increasing or decreasing
from normal demand speed. Under certain failure conditions, % TRQ, % RPM, and Ng SPEED may not be indicating and the possibility of the ENG OUT warning light
and audio activating exists. The most reliable indication of
engine power will be TGT TEMP.

b. Cabin door window jettison. To provide emergency
exit from the cabin, two jettisonable windows are installed
in each cabin door. To release the windows, a handle (under a jettison lever guard) marked EMERGENCY EXIT
PULL AFT, (left side; right side, PULL FWD) on the
inside of the cabin door (Figure 9-1), is moved in the direction of the arrow, releasing the windows. The windows
can then be pushed out.
9.6 EMERGENCY EQUIPMENT (PORTABLE).

f. The term EMER APU START is defined as APU
start to accomplish an emergency procedure.
1. FUEL PUMP switch - APU BOOST.
2. APU CONTR switch - ON.

Emergency equipment consists of two hand held fire extinguishers, one crash ax, and three first aid kits, as shown
in Figure 9-1.
9.7 ENGINE MALFUNCTION - PARTIAL OR COMPLETE POWER LOSS.

9.4 AFTER EMERGENCY ACTION.
After a malfunction of equipment has occurred, appropriate emergency actions have been taken and the helicopter is on the ground, an entry shall be made in the Remarks
Section of DA Form 2408-13-1 describing the malfunction.
Ground and flight operations shall be discontinued until
corrective action has been taken.
9.5 EMERGENCY EXITS.
Emergency exits are shown in Figure 9-1. Emergency
exit release handles are yellow and black striped.

WARNING
For helicopters without a roll-trim actuator, the cyclic shall be held at all times
with the rotor turning. In cases where
emergency exit is required prior to rotor
coasting to a stop, make sure that the cyclic stick is centered until the last crewmember can depart the cockpit. Since the
main rotor shaft has a 3° forward tilt, an
exit to the right rear or left rear will provide the greatest rotor clearance safety.
a. Each cockpit door is equipped with a jettison system
for emergency release of the door assembly. Jettison is done
by pulling a handle marked EMERGENCY EXIT PULL,
on the inside of the door (Figure 9-1). To release the door,
the jettison handle is pulled to the rear; the door may then
be jettisoned by kicking the lower forward corner of the
door. On helicopters equipped with jettisonable cockpit
door windows if the door fails to jettison, the windows may
be removed by pulling the emergency strap inward.

9-2

Change 8

WARNING
Prior to movement of either powercontrol lever, it is imperative that the
malfunctioning engine and the corresponding power-control lever be identified. If the decision is made to shut down
an engine, take at least five full seconds
while retarding the ENG POWER CONT
lever from FLY to IDLE, monitoring %
TRQ, Ng SPEED, TGT TEMP, % RPM,
and ENG OUT warning light on.
The various conditions under which engine failure may
occur, prevent a standard procedure. A thorough knowledge of emergency procedures and flight characteristics will
enable the pilot to respond correctly and automatically in
an emergency. The engine instruments often provide ample
warning of a malfunction before actual engine failure. The
indications of engine malfunction, either partial or complete power loss, may be as follows: Changes in affected
engine % RPM, TGT TEMP, Ng SPEED, % TRQ, ENG
OIL PRESS, % RPM R, LOW ROTOR RPM and/or
ENG OUT warning lights and audio, and change in engine
noise. The amount of change in each depends upon the type
of failure, e.g., compressor stall, as opposed to complete
power loss on one or both engines.
9.8 FLIGHT CHARACTERISTICS.
DUAL-ENGINE FAILURE: The flight characteristics
and the required crewmember control responses after a
dual-engine failure are similar to those during a normal
power-on descent. Full control of the helicopter can be
maintained during autorotational descent. In autorotation,

TM 1-1520-237-10

as airspeed increases above 70 - 80 KIAS, the rate of descent and glide distance increase significantly. As airspeed
decreases below 64 KIAS, the rate of descent will increase
and glide distance will decrease.
SINGLE-ENGINE FAILURE: When one engine has
failed, the helicopter can often maintain altitude and airspeed until a suitable landing site can be selected. Whether
or not this is possible becomes a function of such combined
variables as aircraft weight, density altitude, height above
ground, airspeed, phase of flight, single engine capability,
and environmental response time and control technique
may be additional factors. In addition, these factors should
be taken into consideration should the functioning engine
fail and a dual-engine failure results.
9.9 SINGLE-ENGINE FAILURE - GENERAL.

9.10 SINGLE-ENGINE FAILURE.

WARNING
Do not respond to ENG OUT warning
light and audio until checking TGT
TEMP, Ng SPEED, and % RPM 1 and 2.
1. Collective - Adjust to maintain % RPM R.
2. External cargo/stores - Jettison (if required).

If continued flight is not possible:
3. LAND AS SOON AS POSSIBLE.
If continued flight is possible:

WARNING
4. Establish single-engine airspeed.
When the power available during single
engine operation is marginal or less, consideration should be given to jettisoning
the external stores. The engine anti-ice
and cabin heater switches should be
turned off as necessary to ensure maximum power is available on the remaining
engine.
Crewmember recognition of a single-engine failure and
subsequent action are essential and should be based on the
following general guidelines. At low altitude and low airspeed, it may be necessary to lower the collective only
enough to maintain % RPM R (normal range). At higher
altitude, however, the collective may be lowered significantly to increase % RPM R to 100 percent. When hovering in ground effect, the collective should be used only as
required to cushion the landing, and the primary consideration is in maintaining a level attitude. In forward flight at
low altitude (as in takeoff), when a single-engine capability
to maintain altitude does not exist, a decelerating attitude
will initially be required to prepare for landing. Conversely,
if airspeed is low and altitude sufficient, the helicopter
should be placed in an accelerating attitude to gain sufficient airspeed for single-engine fly away to a selected landing site. The light regions in the height velocity avoid region diagram s (Figures 9-2 and 9-3) define the ground
speed and wheel-height combinations that will permit a safe
landing in the event of an engine failure for various gross
weights at both sea level 15°C (59°F), and 4,000 feet/35°C
(95°F), ambient condition.

5. LAND AS SOON AS PRACTICABLE.
9.11 ENGINE RESTART DURING FLIGHT.
After an engine failure in flight, an engine restart may be
attempted. If it can be determined that it is reasonably safe
to attempt a start, the APU should be used. Use of a crossbleed start could result in a power loss of up to 18% on the
operational engine.
9.12 DUAL-ENGINE FAILURE - GENERAL.
a. If both engines fail, immediate action is required to
make a safe autorotative descent. The altitude and airspeed
(Figure 9-4) at which a two-engine failure occurs will dictate the action to be taken. After the failure, main rotor rpm
will decay rapidly and the aircraft will yaw to the left.
Unless a two-engine failure occurs near the ground, it is
mandatory that autorotation be established immediately.
During cruise, reduce collective immediately to regain %
RPM R and then adjust as required to maintain % RPM
within power off rotor speed limits. The cyclic should be
adjusted as necessary to attain and maintain the desired
airspeed. The recommended airspeed for autorotation is 80
KIAS. Autorotation below 80 knots is not recommended
because the deceleration does not effectively arrest the rate
of descent. Adjusting the cyclic and collective control to
maintain 100% RPM R and 110 KIAS (100 KIAS high
drag) will result in achieving the maximum glide distance.
A landing area must be selected immediately after both
engines fail. Throughout the descent, adjust collective as

Change 8

9-3

TM 1-1520-237-10

A
FIRE
EXTINGUISHER
(COCKPIT)

FIRE
EXTINGUISHER
(CABIN)
PILOT’S
SEAT

FIRST
AID KIT

F
COPILOT’S
SEAT

E
FIRST AID
KIT (CABIN)

BATTERY
CRASH AX
(CABIN)

FIRST
AID KIT

VIEW LOOKING FORWARD
B
BE
ST ION
MU SIT
LE PO
ND E"
G
HA LOS
SIN
LO
"C
IN RE C
O
F
BE
OR
DO

ED
CK
LO
E
OS
CL

EN
OP

HANDLE MUST BE
IN "CLOSE" POSITION
BEFORE CLOSING
DOOR

CABIN AND COCKPIT
DOORS EXTERIOR HANDLE
CABIN DOOR
(SAME FOR RIGHT SIDE)

Figure 9-1. Emergency Exits and Emergency Equipment Diagram (Sheet 1 of 3)

9-4

Change 2

AA0533_1A
SA

TM 1-1520-237-10

DOOR
HANDLE

LATCH LEVER

HANDLE MUST BE
IN "CLOSE" POSITION
BEFORE CLOSING
DOOR

FWD

CABIN DOOR
INTERIOR RELEASE HANDLE
OPEN
POSITION

VIEW LOOKING OUTBOARD
LEFT SIDE
(SAME FOR RIGHT SIDE)
DOOR
HANDLE

KEY SLOT

UNLOCKED
POSITION

JETTISON
LEVER
GUARD

EMERGENCY EXIT
PULL
AFT

LOCKED
POSITION

FWD

CABIN DOOR WINDOW JETTISON LEVER
VIEW LOOKING OUTBOARD
LEFT SIDE
(RIGHT SIDE PULL
FWD)

EMERG EXIST DOOR USE ONLY
WHEN SHIPS BLADES ARE AT A FULL STOP.
OTHERWISE YOU COULD LOOSE YOUR HEAD.
SFDGH 6YUOIP 890PO0P

LOCKED
CLOSE

OPEN

FWD

CABIN DOOR
EXTERIOR RELEASE HANDLE
VIEW LOOKING INBOARD
LEFT SIDE
(SAME FOR RIGHT SIDE)
AA0533_2A
SA

Figure 9-1. Emergency Exits and Emergency Equipment Diagram (Sheet 2 of 3)
necessary to maintain % RPM R within normal range.
Figure 5-1 shows the rotor limitations. % RPM R should
be maintained at or slightly above 100 percent to allow
ample rpm before touchdown.
b. Main rotor rpm will increase momentarily when the
cyclic is moved aft with no change in collective pitch setting. An autorotative rpm of approximately 100 percent
provides for a good rate of descent. % RPM R above 100
percent will result in a higher rate of descent. At 50 to 75
feet AGL, use aft cyclic to decelerate. This reduces airspeed and rate of descent and causes an increase in %
RPM R. The degree of increase depends upon the amount
and rate of deceleration. An increase in % RPM R can be
desirable in that more inertial energy in the rotor system

will be available to cushion the landing. Ground contact
should be made with some forward speed. Pitch attitudes
up to 25° at the point of touchdown normally result in an
adequate deceleration and safe landing. If a rough area is
selected, a steeper deceleration and a touchdown speed as
close to zero as possible should be used. With pitch attitude
beyond 25° there is the possibility of ground contact with
the stabilator trailing edge. It is possible that during the
autorotative approach, the situation may require additional
deceleration. In that case, it is necessary to assume a landing attitude at a higher altitude than normal. Should both
engines fail at low airspeed, initial collective reduction may
vary widely. The objective is to reduce collective as necessary to maintain % RPM R within normal range. In some

Change 8

9-5

TM 1-1520-237-10

WINDOW
EMERGENCY
EXIT STRAP

JETTISON
LEVER
JETTISON
LEVER
KEY SLOT
KEY SLOT

E
OS

CK
LO

EXTERIOR
RELEASE
HANDLE

G
ER
EM

LL
PU
IT
EX

N
LL
WHE A FU
LY
LL
TO
FU
E ON ME
A
US
CO
TO
LL
ES
IST,
FU
A
AD COME
EX
G
TO
R BL ES
ER
AD COME
EM OPTE
R BL ES
LIC
AD
HE OPTE
BL
LIC TER
HE
OP

ED
CK

E
OS
CL

N
LL
WHE A FU
LY
LL
TO
FU
E ON ME
A
US
CO
TO
LL
ES
IST,
A FU
AD COME
EX
G
BL
TO
R
ES
ER
AD COME
EM OPTE
BL
ES
LIC TER
AD
OP
BL
LIC TER
HE
OP
LIC
HE

EXTERIOR
RELEASE
HANDLE

LIC
HE

C
G
(IF INSTALLED)

CABIN DOOR
INTERIOR JETTISON LEVER
VIEW LOOKING INBOARD
LEFT SIDE
(SAME FOR RIGHT SIDE)

F
G
FWD

HANDLE MUST BE
IN "CLOSE"POSITION
BEFORE CLOSING
DOOR

OPEN

CLOSE
LOCKED

INTERIOR COCKPIT DOOR RELEASE
HANDLE
VIEW LOOKING OUTBOARD
LEFT SIDE
(SAME FOR RIGHT SIDE)

AA0533_3
SA

Figure 9-1. Emergency Exits and Emergency Equipment Diagram (Sheet 3 of 3)

9-6

Change 1

TM 1-1520-237-10

instances at low altitude or low airspeed, settling may be so
rapid that little can be done to avoid a hard-impact landing.
In that case, it is critical to maintain a level landing attitude.
Cushion the landing with remaining collective as helicopter
settles to the ground. At slow airspeeds, where altitude permits, apply forward cyclic as necessary to increase airspeed
to about 80 KIAS. Jettison external cargo and stores as
soon as possible to reduce weight and drag, improve autorotational performance, and reduce the chance of damage
to the helicopter on landing.
9.13 DUAL-ENGINE FAILURE.

WARNING
Do not respond to ENG OUT warning
lights and audio until checking TGT
TEMP and % RPM R.
AUTOROTATE.

If %TRQ 1 and 2 are equal, attempt to increase %RPM R with RPM trim switch.
1. Collective - Adjust to control % RPM R.
2. ENG POWER CONT lever - LOCKOUT low
% TRQ/TGT TEMP engine. Maintain %
TRQ approximately 10% below other engine.

3. LAND AS SOON AS PRACTICABLE.
9.15 INCREASING % RPM R.
% RPM R increasing will result from an engine control
system failing to the high side. % RPM 1 and 2 (Np) will
increase with the rotor % RPM R. Increasing the collective
will probably increase the malfunctioning engine’s TGT
TEMP above 900°C. If an engine control unit fails to the
high side:
1. ENG POWER CONT lever - Retard high %
TRQ/TGT TEMP engine, maintain % TRQ
approximately 10% below other engine.

9.14 DECREASING % RPM R.
2. LAND AS SOON AS PRACTICABLE.
If an engine control unit fails to the low side and the
other engine is unable to provide sufficient torque, % RPM
R will decrease.

If the affected engine does not respond to ENG
POWER CONT lever movement in the range between
FLY and IDLE, the HMU may be malfunctioning internally.

CAUTION

If this occurs:
When engine is controlled with engine
power-control lever in lockout, engine response is much faster and the TGT limiting system is inoperative. Care must be
taken not to exceed TGT limits and keeping % RPM R and % RPM 1 and 2 in
operating range.

3. Establish single engine airspeed.
4. Perform EMER ENG SHUTDOWN (affected
engine).

NOTE
If %RPM R reduces from 100% to 95-96%
during steady flight, check %TRQ 1 and 2.

Change 8

9-6.1/(9-6.2 Blank)

TM 1-1520-237-10

UH−60A/EH−60A
HEIGHT VELOCITY AVOID REGIONS
SINGLE−ENGINE FAILURE

EXAMPLE

SEA LEVEL STANDARD

WANTED:

650

A TAKEOFF PROFILE WHICH WILL
PERMIT A SAFE LANDING AFTER
AN ENGINE SUDDENLY BECOMES
INOPERATIVE.

600
550
AVOID AREA

WHEEL HEIGHT ~ FT

500

KNOWN:
AIRCRAFT GROSS WEIGHT = 22,000 LBS
AMBIENT CONDITIONS:
TEMPERATURE = 15oC
PRESSURE ALTITUDE = SEA LEVEL
WIND = 0 KTS

METHOD:
TRACE ALONG GROSS WEIGHT LINE
NOTING WHEEL HEIGHT / AIRSPEED
COMBINATIONS WHICH WILL KEEP
THE TAKEOFF PROFILE BELOW AND
TO THE RIGHT OF THE AVOID REGION.

450
400
350

GROSS
WEIGHT
~ 1000 LBS

300
250
200
150

16,000
LB

100

18,000
LB

50
POINT
A
B
C
D

AIRSPEED
10
20
30
42

WHEEL HEIGHT
10
10
15
155

20,000
LB

D
22,000
LB

24,500
LB

0
0

AIRSPEED ~ KTS

4000 FT 35 OC (95 oF)
850
800
750
AVOID AREA

700
650

WHEEL HEIGHT ~ FT

600
550
500
450

GROSS
WEIGHT
~ 1000 LBS

400
350
300

22,000
LB

250
200
150

NOTE:
BASED ON AN ETF OF .85 OF
MAXIMUM RATED POWER.
WEIGHTS GREATER THAN
22,000 LBS ARE FOR FERRY
MISSION ONLY, FOR WHICH A
FLIGHT RELEASE IS REQUIRED.

14,000
LB

12,000
LB

100

16,000
LB

18,000
LB

20,000
LB

50
0
0

AIRSPEED ~ KTS

DATA BASIS: CALCULATED

Figure 9-2. Height Velocity Diagram

UH−60A

70
AA0673C
SA

9-7

TM 1-1520-237-10

UH−60L
HEIGHT VELOCITY AVOID REGIONS
SINGLE ENGINE FAILURE

EXAMPLE

SEA LEVEL STANDARD
500

WANTED

KNOWN
AIRCRAFT GROSS WEIGHT = 22,000 LBS
AMBIENT CONDITIONS:
TEMPERATURE = 15oC
PRESSURE ALTITUDE = SEA LEVEL
WIND = 0 KTS

400

WHEEL HEIGHT ~ FT

A TAKEOFF PROFILE WHICH WILL
PERMIT A SAFE LANDING AFTER
AN ENGINE SUDDENLY BECOMES
INOPERATIVE.

24,500 LB
300

22,000 LB

AVOID AREA

20,000 LB
200
18,000 LB
C

METHOD
100
TRACE ALONG GROSS WEIGHT LINE
NOTING WHEEL HEIGHT / AIRSPEED
COMBINATIONS WHICH WILL KEEP
THE TAKEOFF PROFILE BELOW AND
TO THE RIGHT OF THE AVOID REGION.
POINT
A
B
C

AIRSPEED
10
20
30

B
A
0
0

WHEEL HEIGHT
21
33
150

AIRSPEED ~ KTS

4000 FT 35 oC (95 oF)
800

700

WHEEL HEIGHT ~ FT

600

AVOID AREA

22,000 LB

500

20,000 LB
400
18,000 LB
300

200
16,000 LB

100

NOTE
BASED ON AN ETF OF .90 OF MAXIMUM
RATED POWER.

DATA BASIS: ESTIMATED

0
0

AA1639A
SA

Figure 9-3. Height Velocity Diagram

9-8

AIRSPEED ~ KTS

UH−60L

TM 1-1520-237-10

AUTOROTATION
CLEAN CONFIGURATION
100% RPM R ZERO WIND

GLIDE RATIO ~

NAUTICAL MILES PER 1000 FT

.75
.70
.65

A GR VARIATION
OF .05 MAY
RESULT FROM
FLIGHT CONDITION
ABOVE / BELOW
THE NOMINAL

.60
AIRSPEED FOR
MAXIMUM GLIDE
110 KIAS

.55
.50

A R / D VARIATION
OF 300 FT / MIN
MAY RESULT
FROM FLIGHT
CONDITIONS
ABOVE / BELOW
THE NOMINAL

4600
4400
4200
MAXIMUM
AUTOROTATION
AIRSPEED FOR
GW’S ABOVE
16825 LB

RATE OF DESCENT ~ FT/MIN

4000
3800
3600

MAXIMUM
AIRSPEED
FOR GW’S
UP TO
16825 LB

3400
3200
RECOMMENDED
AUTOROTATIONAL
AIRSPEED = 80 KIAS

3000
GROSS
WEIGHT
~ 1000 LB

2800

2600

20
18

2400

16
2200

14
12

2000
30

100

110

120

130

140

150

INDICATED AIRSPEED ~ KNOTS

EXAMPLE
WANTED:

KNOWN:

METHOD:

RATE OF DESCENT AND GLIDE RATIO IN STEADY
STATE AUTOROTATION AT IAS’S OF:
A. 130 KIAS
B. 50 KIAS

RPM R = 100%
GROSS WEIGHT = 18,000 LB

A. ENTER AUTOROTATION CHART AT 130 KIAS,
MOVE UP TO INTERSECTION OF R / D LINE,
MOVE LEFT, READ RD = 3,560 FT / MIN
B. ENTER AUTOROTATION CHART AT 50 KIAS,
MOVE UP TO 18,000 LB GW LINE, MOVE LEFT,
READ R / D = 2,325 FT / MIN

DATA BASIS: FLIGHT TEST

THE GLIDE RATIO IS NOT PROVIDED FOR LOW
AIRSPEEDS

AA0322B
SA

Figure 9-4. Autorotative Glide Distance Chart
9-9

TM 1-1520-237-10

AUTOROTATION
HIGH DRAG CONFIGURATION
100% RPM R ZERO WIND
NOTE: DASH LINE FOR FERRY
MISSION ONLY

NAUTICAL MILES TRAVELED

PER 1000 FT ALTITUDE LOSS

GLIDE RATIO ~ GR

.75

.70

.65

.60

.55
AIRSPEED FOR
MAXIMUM GLIDE
100 KIAS

.50

.45

4600

4400

4200

RATE OF DESCENT ~ FT / MIN

4000
MAXIMUM
AIRSPEED
FOR GW’S
UP TO
16,825 LB

3800

3600

3400

3200

RECOMMENDED
AUTOROTATIONAL
AIRSPEED =
80 KIAS

GROSS WEIGHT
= 1000 LB
3000

MAXIMUM
AUTOROTATION
AIRSPEED FOR
GW’S ABOVE
16,825 LB

24
2800
22
2600
20
2400

18
16
14

2200

2000
40

DATA BASIS: FLIGHT TEST

100

120

140

INDICATED AIRSPEED ~ KNOTS

AA1253A
SA

Figure 9-5. Autorotative Glide Distance Chart-High Drag
9-10

TM 1-1520-237-10

5. Refer to single engine failure emergency procedure.
9.16 % RPM INCREASING/DECREASING (OSCILLATION).
It is possible for a malfunction to occur that can cause
the affected engine to oscillate. The other engine will respond to the change in power by also oscillating, usually
with smaller amplitudes. The engine oscillations will cause
torque oscillations. The suggested pilot corrective action is
to pull back the ENG POWER CONT lever of the suspected engine until oscillation stops. If the oscillation continues, the ENG POWER CONT lever should be returned
to FLY position and the other ENG POWER CONT lever
pulled back until the oscillation ceases. Once the malfunctioning engine has been identified, it should be placed in
LOCKOUT and controlled manually.
1. Slowly retard the ENG POWER CONT lever
on the suspected engine.

If the oscillation stops:
2. Place that engine in LOCKOUT and manually
control the power.

3. LAND AS SOON AS PRACTICABLE.
If the oscillation continues:
4. Place the ENG POWER CONT lever back to
FLY and retard the ENG POWER CONT lever of the other engine.

When the oscillation stops:
5. Place the engine in LOCKOUT, manually control the power.

6. LAND AS SOON AS PRACTICABLE.
9.17 % TRQ SPLIT BETWEEN ENGINES 1 AND 2.
It is possible for a malfunction to occur that can cause a
% TRQ split between engines without a significant change
in % RPM R. The % TRQ split can be corrected by manual
control of the ENG POWER CONT lever on the affected
engine.
1. If TGT TEMP of one engine exceeds the limiter ( 700 849°C , 701C 872°C with low power
engine above 50% TRQ or 896°C with low
power engine below 50% TRQ), retard ENG

POWER CONT lever on that engine to reduce
TGT TEMP. Retard the ENG POWER
CONT lever to maintain torque of the manually controlled engine at approximately 10%
below the other engine.
2. If TGT TEMP limit on either engine is not
exceeded, slowly retard ENG POWER CONT
lever on high % TRQ engine and observe %
TRQ of low power engine.
3. If % TRQ of low power engine increases,
ENG POWER CONT lever on high power engine - Retard to maintain % TRQ approximately 10% below other engine (The high
power engine has been identified as a high side
failure).
4. If % TRQ of low power engine does not increase, or % RPM R decreases, ENG
POWER CONT lever - Return high power engine to FLY (The low power engine has been
identified as a low side failure).
5. If additional power is required, low power
ENG POWER CONT lever, momentarily
move to LOCKOUT and adjust to set % TRQ
approximately 10% below the other engine.

6. LAND AS SOON AS PRACTICABLE.
9.18 ENGINE COMPRESSOR STALL.
An engine compressor stall is normally recognized by a
noticeable bang or popping noise and possible aircraft yaw.
These responses are normally accompanied by the rapid
increase in TGT TEMP and fluctuations in Ng SPEED, %
TRQ, and % RPM reading for the affected engine. In the
event of a compressor stall:
1. Collective - Reduce.
If condition persists:
2. ENG POWER CONT lever (affected engine) Retard (TGT TEMP should decrease.)
3. ENG POWER CONT lever (affected engine) FLY.

If stall condition recurs:
4. EMER ENG SHUTDOWN (affected engine).

Change 8

9-11

TM 1-1520-237-10

5. Refer to single-engine failure emergency procedure.

9.19 ENGINE OIL FILTER BYPASS CAUTION
LIGHT ON, ENGINE CHIP CAUTION LIGHT ON,
ENG OIL PRESS HIGH/LOW, ENGINE OIL TEMP
HIGH, ENGINE OIL TEMP CAUTION LIGHT ON, ENGINE OIL PRESS CAUTION LIGHT ON.

practical, the pilot should reduce speed to 80 KIAS. This
will reduce the criticality of having exactly correct rotor
speed 100%.
1. ENG POWER CONT levers - Adjust as required to control % RPM.

2. LAND AS SOON AS POSSIBLE.

1. ENG POWER CONT lever - Retard to reduce
torque on affected engine.

9.22 ROTORS, TRANSMISSIONS AND DRIVE SYSTEMS.

If oil pressure is below minimum limits or if oil temperature remains above maximum limits:

9.22.1 Loss of Tail Rotor Thrust. Failure of the tail
rotor gearbox, intermediate gearbox or tail rotor drive shaft
will result in a loss of tail rotor thrust. The nose of the
helicopter will yaw right regardless of the airspeed at which
the failure occurs. Continued level flight may not be possible following this type failure. Loss of tail rotor thrust at
low speed will result in rapid right yaw. At higher airspeed,
right yaw may develop more slowly but will continue to
increase. Autorotation should be entered promptly. Retard
ENG POWER CONT levers to OFF position during deceleration. Every effort should be made to establish and
maintain an autorotative glide at or above 80 KIAS. This
will maximize the effectiveness of the deceleration during
the landing sequence. If autorotation entry is delayed, large
sideslip angles can develop causing low indicated airspeed
with the stabilator programming down. This can make it
more difficult to establish or maintain adequate autorotative
airspeed.

2. EMER ENG SHUTDOWN (affected engine).
3. Refer to single-engine failure emergency procedure.

9.20 ENGINE HIGH-SPEED SHAFT FAILURE.
Failure of the shaft may be complete or partial. A partial
failure may be characterized at first by nothing more than a
loud high-speed rattle and vibration coming from the engine area. A complete failure will be accompanied by a
loud bang that will result in a sudden % TRQ decrease to
zero on the affected engine. % RPM of affected engine
will increase until overspeed system is activated.
1. Collective - Adjust.
2. EMER ENG SHUTDOWN (affected engine).
Do not attempt to restart.
3. Refer to single-engine failure emergency procedure.

9.21 LIGHTNING STRIKE.

WARNING
Lightning strikes may result in loss of automatic flight control functions, engine
controls, and/or electric power.
Lightning strike may cause one or both engines to immediately produce maximum power with no TGT limiting
or overspeed protection. Systems instruments may also be
inoperative. If this occurs, the flight crew would have to
adjust to the malfunctioning engine(s) power-control lever(s) as required to control % RPM by sound and feel. If

9-12

Change 9

1. AUTOROTATE.
2. ENG POWER CONT levers - OFF (when intended point of landing is assured).

9.22.2 Loss of Tail Rotor Thrust at Low Airspeed/
Hover.
Loss of tail rotor thrust at slow speed may result in
extreme yaw angles and uncontrolled rotation to the right.
Immediate collective pitch reduction should be initiated to
reduce the yaw and begin a controlled rate of descent. If the
helicopter is high enough above the ground, initiate a
power-on descent. Collective should be adjusted so that an
acceptable compromise between rate of turn and rate of
descent is maintained. At approximately 5 to 10 feet above
touchdown, initiate a hovering autorotation by moving the
ENG POWER CONT levers - OFF.
1. Collective - Reduce.
2. ENG POWER CONT levers - OFF (5 to 10
feet above touchdown).

TM 1-1520-237-10

9.22.3 TAIL ROTOR QUADRANT Caution Light On
With No Loss of Tail Rotor Control.

WARNING
If the helicopter is shut down and/or hydraulic power is removed with one tail rotor cable failure, disconnection of the
other tail rotor cable will occur when
force from the boost servo cannot react
against control cable quadrant spring tension. The quadrant spring will displace
the cable and boost servo piston enough
to unlatch the quadrant cable.

malfunction can produce about 30 pounds at the pedal. An
internally jammed yaw trim actuator can produce up to 80
pounds until clutch slippage relieves this force. The pilot
can override any yaw trim force by applying opposite pedal
firmly and then turning off trim. A malfunction within the
yaw boost servo or tail rotor servo can produce much higher
force at the pedals and the affected servo must be turned
off. Hardover failure of the yaw boost servo will increase
control forces as much as 250 pounds on the pedals.
1. Apply pedal force to oppose the drive.
2. TRIM switch - Off.

If normal control forces are not restored:
3. BOOST switch - Off.

Loss of one tail rotor cable will be indicated by illumination of TAIL ROTOR QUADRANT caution light. No
change in handling characteristics should occur.
LAND AS SOON AS PRACTICABLE.

If control forces, normal for boost off flight are not restored:
4. BOOST switch - ON.

9.22.4 TAIL ROTOR QUADRANT Caution Light On
With Loss of Tail Rotor Control.

5. TAIL SERVO switch - BACKUP, if tail rotor
is not restored.

a. If both tail rotor control cables fail, a centering spring
will position the tail rotor servo linkage to provide 10-1/2
degrees of pitch. This will allow trimmed flight at about 25
KIAS and 145 KIAS (these speeds will vary with gross
weight). At airspeed below 25 and above 145 KIAS, right
yaw can be controlled by reducing collective. Between 25
and 145 KIAS, left yaw can be controlled by increasing
collective.

a. If the tail rotor quadrant becomes jammed,
collective control is available, except that
low collective with right pedal or high collective with a left pedal will be restricted.
With a quadrant jam, complete collective
travel is available for most control combinations, provided the pedals are allowed to
move as the collective is displaced.

b. A shallow approach to a roll-on landing technique is
recommended. During the approach, a yaw to the left will
occur. As the touchdown point is approached, a mild deceleration should be executed to reduce airspeed. As collective
is increased to cushion touchdown, the nose of the helicopter will yaw right. Careful adjustment of collective and deceleration should allow a tail-low touchdown with approximate runway alignment. Upon touchdown, lower collective
carefully. Use brakes to control heading.

b. If tail rotor pitch becomes fixed during decreased power situations (right pedal applied), the nose of the helicopter will turn
to the right when power is applied, possibly even greater than complete loss of tail
rotor thrust. Some conditions may require
entry into autorotation to control yaw rate.
If continued flight is possible, a shallow
approach at about 80 KIAS to a roll-on
landing should be made. As the touchdown
point is approached, a mild deceleration
should be executed at about 15 to 25 feet
to reduce airspeed to about 40 KIAS. As
collective is increased to cushion touchdown, the nose of the helicopter will turn
to the right. Careful adjustment of collective and deceleration should allow a taillow touchdown with approximate runway
alignment. Upon

1. Collective - Adjust.
2. LAND AS SOON AS PRACTICABLE.
9.22.5 Pedal Bind/Restriction or Drive With No Accompanying Caution Light. If pedal binding, restriction, or driving occurs with no caution light the cause may
not be apparent. A Stability Augmentation System/Flight
Path Stabilization (SAS/FPS) computer induced yaw trim

Change 8

9-13

TM 1-1520-237-10

touchdown, lower collective carefully and
use brakes to control heading.
c. If tail rotor pitch becomes fixed during increased power situations (left pedal applied), the nose of the helicopter will turn
left when collective is decreased. Under
these conditions, powered flight to a prepared landing site and a powered landing is
possible since the sideslip angle will probably be corrected when power is applied
for touchdown. Adjust approach speed and
rate of descent to maintain a sideslip angle
of less than 20°. Sideslip angle may be reduced by either increasing airspeed or collective. Execute a decelerated touchdown
tailwheel first, and cushion landing with
collective. Upon touchdown, lower collective carefully and use brakes to control
heading.
6. LAND AS SOON AS PRACTICABLE.
9.22.6 #1 TAIL RTR SERVO Caution Light On and
BACK-UP PUMP ON Advisory Light Off or #2 TAIL
RTR SERVO ON Advisory Light Off. Automatic
switch-over did not take place.

9.22.8 CHIP INPUT MDL LH or RH Caution Light
On.
1. ENG POWER CONT lever on affected engine
- IDLE.

2. LAND AS SOON AS POSSIBLE.
9.22.9 CHIP MAIN MDL SUMP, CHIP ACCESS MDL
LH or RH, CHIP TAIL XMSN or CHIP INT XMSN/
TAIL XMSN OIL TEMP or INT XMSN OIL TEMP Caution Light On.
LAND AS SOON AS POSSIBLE.
9.22.10 MAIN TRANSMISSION FAILURE.

WARNING
If % RPM R decreases from 100% to below 96% with an increase in torque during steady flight with no engine malfunction, the main transmission planetary
carrier may have failed. During a main
transmission planetary carrier failure, it
may be impossible to maintain % RPM R
at 100%.

1. TAIL SERVO switch - BACKUP.
2. BACKUP HYD PUMP switch - ON.

3. LAND AS SOON AS PRACTICABLE.
9.22.7 MAIN XMSN OIL PRESS Caution Light On/
XMSN OIL PRESS LOW/XMSN OIL TEMP HIGH or
XMSN OIL TEMP Caution Light On. Loss of cooling
oil supply will lead to electrical and/or mechanical failure
of main generators. If the malfunction is such that oil pressure decays slowly, the generators may fail before MAIN
XMSN OIL PRESS caution light goes on.
1. LAND AS SOON AS POSSIBLE.

NOTE
Decreasing % RPM R may be accompanied
by a drop in transmission oil pressure of 10
PSI or more, and possible unusual helicopter
vibrations.
1. Collective - Adjust only enough to begin a descent with power remaining applied to the main
transmission throughout the descent and landing.
2. LAND AS SOON AS POSSIBLE.
9.23 FIRE.

If time permits:
2. Slow to 80 KIAS.
3. EMER APU START.
4. GENERATORS NO. 1 and NO. 2 switches OFF.

9-14

Change 10

WARNING
If AC electrical power is not available,
only the reserve fire bottle can be discharged and fire extinguishing capability
for the #2 engine will be lost.

TM 1-1520-237-10

The safety of helicopter occupants is the primary consideration when a fire occurs; therefore, it is imperative that
every effort be made to extinguish the fire. On the ground,
it is essential that the engine be shut down, crew and passengers evacuated, and fire fighting begun immediately. If
time permits, a 9Mayday9 radio call should be made before
the electrical power is OFF to expedite assistance from firefighting equipment and personnel. If the helicopter is airborne when a fire occurs, the most important single action
that can be taken by the pilot is to land. Consideration must
be given to jettisoning external stores and turning FUEL
BOOST PUMPS and XFER PUMPS off prior to landing.

9.23.5 Electrical Fire In Flight. Prior to shutting off all
electrical power, the pilot must consider the equipment that
is essential to a particular flight environment which will be
affected, e.g., flight instruments, flight controls, etc. If a
landing cannot be made as soon as possible the affected
circuit may be isolated by selectively turning off electrical
equipment and/or pulling circuit breakers.
1. BATT and GENERATORS switches - OFF.

2. LAND AS SOON AS POSSIBLE.
9.24 SMOKE AND FUME ELIMINATION.

9.23.1 Engine/Fuselage Fire On Ground.
1. ENG POWER CONT levers - OFF.

WARNING

2. ENG EMER OFF handle - Pull if applicable.
3. FIRE EXTGH switch - MAIN/RESERVE as
required.

9.23.2 APU Compartment Fire.
1. APU fire T-handle - Pull.
2. FIRE EXTGH switch - MAIN/RESERVE as
required.

If battery overheats, do not remove battery cover or attempt to disconnect or remove battery. Battery fluid will cause
burns, and an overheated battery could
cause thermal burns and may explode.
Smoke or fumes in the cockpit/cabin can be eliminated
as follows:
1. Airspeed - 80 KIAS or less.

9.23.3 APU OIL TEMP HI Caution Light On.

2. Cabin doors and gunner’s windows - Open.

APU CONTR switch - OFF. Do not attempt restart
until oil level has been checked.

3. Place helicopter out of trim.
4. LAND AS SOON AS PRACTICABLE.

9.23.4 Engine Fire In Flight.
9.25 FUEL SYSTEM.

WARNING
Attempt to visually confirm fire before engine shutdown or discharging extinguishing agent.
1. ENG POWER CONT lever (affected engine) OFF.
2. ENG EMER OFF handle - Pull.
3. FIRE EXTGH switch - MAIN/RESERVE as
required.

4. LAND AS SOON AS POSSIBLE.

9.25.1 #1 or #2 FUEL FLTR BYPASS Caution Light
On.
1. ENG FUEL SYS selector on affected engine XFD.

2. LAND AS SOON AS PRACTICABLE.
9.25.2 #1 and #2 FUEL FLTR BYPASS Caution
Lights On.
LAND AS SOON AS POSSIBLE.
9.25.3 #1 FUEL LOW and #2 FUEL LOW Caution
Lights On.
LAND AS SOON AS PRACTICABLE.

Change 10

9-15

TM 1-1520-237-10

9.25.4 #1 or #2 FUEL PRESS Caution Light On.
a. If the light illuminates, flameout is possible. Do not
make rapid collective movements. This emergency procedure has been written to include corrective action for critical situations. Critical situations are those where the loss of
an engine represents a greater hazard than the possibility of
pressurizing a fuel leak.
If the caution light illuminates and the situation is critical:
1. FUEL BOOST PUMP CONTROL switches NO. 1 PUMP and NO. 2 PUMP - ON.

2. LAND AS SOON AS PRACTICABLE.
b. This portion of the emergency procedure has been
written to provide the best method of isolating the cause of
the failure and prescribing the proper corrective action
when the situation is not critical. This portion of the emergency procedure assumes the FUEL BOOST PUMP
CONTROL switches are OFF when the malfunction occurs.
If the situation is not critical:
1. ENG FUEL SYS selector on affected engine XFD.

If caution light stays on:
2. FUEL BOOST PUMP CONTROL switches NO. 1 PUMP and NO. 2 PUMP - ON.

4. GENERATORS NO. 1 and NO. 2 switches OFF.
5. EMER APU START.
6. SAS 1 switch - ON.

7. LAND AS SOON AS PRACTICABLE.
9.26.2 #1 or #2 GEN Caution Light On.

CAUTION

When the #1 ac generator is failed, and
the backup pump circuit breaker is out,
turn off ac electrical power before resetting the backup pump power circuit
breaker, to avoid damaging the current
limiters.
1. Affected GENERATORS switch - RESET;
then ON.

If caution light remains on:
2. Affected GENERATORS switch - OFF.

9.26.3 #1 and #2 CONV Caution Lights On.
1. Unnecessary dc electrical equipment - OFF.

NOTE
If caution light stays on:
3. FUEL BOOST PUMP CONTROL switches NO. 1 PUMP and NO. 2 PUMP - OFF.

4. LAND AS SOON AS PRACTICABLE.
9.26 ELECTRICAL SYSTEM.
9.26.1 #1 and #2 Generator Failure (#1 and #2
CONV and AC ESS BUS OFF Caution Lights On).

When only battery power is available,
NICAD battery life is about 22 minutes day
and 14 minutes night for a battery 80%
charged. SLAB battery life is about 38 minutes day and 24 minutes night for a battery
80% charged.
2. LAND AS SOON AS PRACTICABLE.
9.26.4 BATTERY FAULT Caution Light On.

1. SAS 1 switch - Press off.

2. Airspeed - Adjust (80 KIAS or less).
3. GENERATORS NO. 1 and NO. 2 switches RESET; then ON.

If caution lights remain on:

9-16

Change 10

1. BATT switch - OFF; then ON. If BATTERY
FAULT caution light goes on, cycle BATT
switch no more than two times.

If caution light remains on:
2. BATT switch - OFF.

TM 1-1520-237-10

9.26.5 BATT LOW CHARGE Caution Light On.
If caution light goes on after AC power is applied:
1. BATT switch - OFF; then ON. About 30 minutes may be required to recharge battery.

9.27.5 #1 or #2 PRI SERVO PRESS Caution Light
On. Illumination of #1 or #2 PRI SERVO PRESS caution
light can be caused by inadvertently placing the SVO OFF
switch on either collective control head in 1ST STG or
2ND STG position. Before initiating emergency procedure
action, the pilots should check that both SVO OFF
switches are centered.

If caution light goes on in flight:
2. BATT switch - OFF, to conserve remaining
battery charge.

1. SVO OFF switch - 1ST STG or 2ND STG as
applicable.
2. LAND AS SOON AS POSSIBLE.

9.27 HYDRAULIC SYSTEM.
9.27.1 #1 HYD PUMP Caution Light On.
1. TAIL SERVO switch - BACKUP; then NORMAL.

2. LAND AS SOON AS PRACTICABLE.

9.27.6 #1 RSVR LOW and #1 HYD PUMP Caution
Lights On With BACK-UP PUMP ON Advisory Light
On.
1. LAND AS SOON AS PRACTICABLE.
If the BACK-UP RSVR LOW caution light also goes
on:

9.27.2 #2 HYD PUMP Caution Light On.
2. SVO OFF switch - 1ST STG.
1. POWER ON RESET switches - Simultaneously press, then release.

2. LAND AS SOON AS PRACTICABLE.
9.27.3 #1 and #2 HYD PUMP Caution Lights On.
LAND AS SOON AS POSSIBLE. Restrict
control movement to moderate rates.

WARNING
If #2 PRI SERVO PRESS caution light
goes on, establish landing attitude, minimize control inputs and begin a descent.
3. LAND AS SOON AS POSSIBLE.

9.27.4 #1 or #2 HYD PUMP Caution Light On and
BACK-UP PUMP ON Advisory Light Off. Loss of
both the No. 1 hydraulic pump and backup pump results in
both stages of the tail-rotor servo being unpressurized. The
yaw boost servo is still pressurized and the mechanical control system is still intact allowing limited tail-rotor control.
Because of the limited yaw control range available, a
roll-on landing 40 KIAS or above is required. Loss of both
the No. 2 hydraulic pump and the backup pump results in
the loss of pilot-assist servos.

9.27.7 #2 RSVR LOW and #2 HYD PUMP Caution
Lights On With BACK-UP PUMP ON Advisory Light
On.
1. POWER ON RESET switches - Simultaneously press then release.

2. LAND AS SOON AS PRACTICABLE.
If BACK-UP RSVR LOW caution light also goes on:

1. Airspeed - Adjust to a comfortable airspeed.
3. SVO OFF switch - 2ND STG.
2. BACKUP HYD PUMP switch - ON.

If BACK-UP PUMP ON advisory light remains off:
3. FPS and BOOST switches - Off (for #2 HYD
PUMP caution light).

4. LAND AS SOON AS POSSIBLE.

WARNING
If #1 PRI SERVO PRESS caution light
goes on, establish landing attitude, minimize control inputs, and begin a descent.

Change 10

9-17

TM 1-1520-237-10

4. LAND AS SOON AS POSSIBLE.
9.27.8 #2 RSVR LOW Caution Light On.
Pilot assist servos will be isolated; if they remain isolated, proceed as follows:
1. BOOST and FPS switches - Off.

2. LAND AS SOON AS PRACTICABLE.
NOTE
Because the logic module will close the
valve supplying pressure to the pilot-assist
servos, BOOST SERVO OFF, SAS OFF,
and TRIM FAIL caution lights will be on.
9.27.9 Collective Boost Servo Hardover/Power
Piston Failure. Hardover failure of the collective boost
servo will increase control forces (as much as 150 pounds)
in the collective. The increased control forces can be immediately eliminated by shutting off the boost servo. Resulting control loads will be the same as for in-flight boost
servo off.
1. BOOST switch - Off.

2. LAND AS SOON AS PRACTICABLE.
9.27.10 Pitch Boost Servo Hardover. Hardover failure of the pitch boost servo will increase the longitudinal
cyclic control forces (approximately 20 pounds). The increased control forces can be immediately eliminated by
shutting off SAS.
1. SAS (1 and 2) and FPS switches - Off.

2. LAND AS SOON AS PRACTICABLE.
9.27.11 BOOST SERVO OFF Caution Light On.
Lighting of the BOOST SERVO OFF caution light with
no other caution lights on indicates a pilot valve jam in
either the collective or yaw boost servo. Control forces in
the affected axis will be similar to flight with boost off.
1. BOOST switch - Off.

2. LAND AS SOON AS PRACTICABLE.

9-18

Change 10

9.28 LANDING AND DITCHING.
9.28.1 Emergency Landing In Wooded Areas.
Power Off.
1. AUTOROTATE. Decelerate helicopter to stop
all forward speed at treetop level.
2. Collective adjust to maximum before main rotor contacts tree branches.
9.28.2 Ditching - Power On. The decision to ditch the
helicopter shall be made by the pilot when an emergency
makes further flight unsafe.
1. Approach to a hover.
2. Cockpit doors jettison and cabin doors open
prior to entering water.
3. Pilot shoulder harness - Lock.
4. Survival gear - Deploy.
5. Personnel, except pilot, exit helicopter.
6. Fly helicopter downwind a safe distance and
hover.
7. ENG POWER CONT levers - OFF.
8. Perform hovering autorotation, apply full collective to decay rotor RPM as helicopter settles.
9. Position cyclic in direction of roll.
10. Exit when main rotor has stopped.
9.28.3 Ditching - Power Off. If ditching is imminent,
accomplish engine malfunction emergency procedures.
During descent, open cockpit and cabin doors. Decelerate
to zero forward speed as the helicopter nears the water.
Apply full collective as the helicopter nears the water.
Maintain a level attitude as the helicopter sinks and until it
begins to roll; then apply cyclic in the direction of the roll.
Exit when the main rotor is stopped.
1. AUTOROTATE.

TM 1-1520-237-10

2. Cockpit doors jettison and cabin doors open
prior to entering water.

craft after landing. Exit when main rotor
has stopped.

3. Cyclic - Position in direction of roll.

1. LAND AS SOON AS POSSIBLE.

4. Exit when main rotor has stopped.

2. EMER ENG(S) SHUTDOWN after landing.

9.29 FLIGHT CONTROL/MAIN-ROTOR
MALFUNCTIONS.

SYSTEM

a. Failure of components within the flight control system may be indicated through varying degrees of feedback,
binding, resistance, or sloppiness. These conditions should
not be mistaken for malfunction of the AFCS.
b. Failure of a main rotor component may be indicated
by the sudden onset or steady increase in main rotor vibration or unusual noise. Severe changes in lift characteristics
and/or balance condition can occur due to blade strikes,
skin separation, shift or loss of balance weights or other
material. Malfunctions may result in severe main rotor flapping. The severity of vibrations may be minimized by reducing airspeed.

9.29.1 SAS Failure With No Failure/Advisory Indication. Erratic electrical input to a SAS actuator can result in moderate rotor tip path oscillations that are often
accompanied with pounding sounds or 9knocking9 which
may be felt in the cyclic or pedal controls. No SAS malfunction, however, can physically drive the pilots’ flight
controls. Failure of SAS 2 is usually but not necessarily
followed by a failure/advisory indication. Failure of a SAS
1 component will not be accompanied by a failure/advisory
indication as SAS 1 does not contain diagnostic capabilities.
If the helicopter experiences erratic motion of the rotor
tip path without failure/advisory indication:

If the main rotor system malfunctions:

WARNING
Danger exists that the main rotor system
could collapse or separate from the air-

Change 10

9-18.1/(9-18.2 Blank)

TM 1-1520-237-10

1. SAS 1 switch - Off.

If condition persists:

If the airspeed fault light remains illuminated on the
AFCS panel:

2. SAS 1 switch - ON.

NOTE
3. SAS 2 switch - Off.

If malfunction still persists:
4. SAS 1 and FPS switches - Off.

9.29.2 SAS 2 Failure Advisory Light On.
POWER ON RESET switches - Simultaneously press and then release.

Use of the cyclic mounted stabilator slew-up
switch should be announced to the crew to
minimize cockpit confusion.
3. Manually slew stabilator - Adjust to 0° if above
40 KIAS. The preferred method of manually
slewing the stabilator up is to use the cyclic
mounted stabilator slew-up switch.
4. LAND AS SOON AS PRACTICABLE.

9.29.3 SAS OFF Caution Light On.
9.29.5 Pitch, Roll or Yaw/Trim Hardover.
FPS switch - Off.
9.29.4 FLT PATH STAB Caution Light On.
a. An FPS malfunction will be detected by the SAS/
FPS computer, which will disengage FPS function in the
applicable axis and light the FLT PATH STAB caution
light and corresponding FAILURE ADVISORY light.
b. EH With the Mode Select Panel switch in the IINS/
IINS position, a failure of the IINS gyro will cause a failure
of the FPS and may possibly cause FPS/SAS 2 to become
erratic in roll motion. In addition to the failure indications
on the IINS Control Display screen, the GYRO segment on
the Failure Advisory Panel will illuminate. The copilot’s
VSI will fail with a ATT warning flag and both HSI’s will
fail with HDG warning flags. The aircraft may drift in
pitch, roll and/or yaw axis due to FPS failure.
1.

SYSTEMS SELECT - DG/VG.

2. POWER ON RESET switches - Simultaneously press and then release.

If failure returns, control affected axis manually.

WARNING

a. A pitch FPS/trim hardover will cause a change in
pitch attitude and a corresponding longitudinal cyclic
movement of about 1/2 inch. This condition will be detected by the SAS/FPS computer which will disengage FPS
and trim functions in the pitch axis and light the FLT
PATH STAB and TRIM FAIL caution lights.
b. A roll FPS/trim hardover will be characterized by a
1/2 inch lateral stick displacement, resulting in a corresponding roll rate and a constant heading sideslip condition, caused by the yaw FPS attempting to maintain heading. The SAS/FPS computer will detect the hardover
condition and disengage lateral trim and illuminate the FLT
PATH STAB and TRIM FAIL caution lights.
c. A yaw FPS/trim hardover is characterized by an improper motion of the pedals, resulting in about 1/4 inch of
pedal motion followed by a corresponding change in helicopter heading trim. This condition will be detected by the
SAS/FPS computer, which will disengage trim and FPS
functions in the yaw axis and light the FLT PATH STAB
and TRIM FAIL caution lights.
If failure occurs:
POWER ON RESET switches - Simultaneously press and then release.
If failure returns, control affected axis manually.

If the airspeed fault advisory light is illuminated, continued flight above 70 KIAS
with the stabilator in the AUTO MODE is
unsafe since a loss of the airspeed signal
from the remaining airspeed sensor would
result in the stabilator slewing full-down.

9.29.6 Trim Actuator Jammed. Both yaw and roll trim
actuators incorporate slip clutches to allow pilot and copilot
inputs if either actuator should jam. The forces required to
override the clutches are 80 pounds maximum in yaw and
13 pounds maximum in roll.

Change 7

9-19

TM 1-1520-237-10

LAND AS SOON AS PRACTICABLE.
9.30 STABILATOR MALFUNCTION - AUTO MODE
FAILURE.
An Auto Mode Failure will normally result in the stabilator failing in place. The indications to the pilots of the
failure are a beeping audio warning, and MASTER CAUTION and STABILATOR caution lights illuminating
when the automatic mode fails. The position of failure may
vary from the ideal programmed position by 10° at 30
KIAS to 4° at 150 KIAS. If an approach is made with the
stabilator fixed 0°, the pitch attitude may be 4° to 5° higher
than normal in the 20 to 40 KIAS range.

1. Cyclic mounted stabilator slew-up switch - Adjust if necessary to arrest or prevent nose down
pitch rate.
2. AUTO CONTROL switch - Press ON once
after establishing a comfortable airspeed.

If automatic control is not regained:
3. Manually slew stabilator - Adjust to 0° for
flight above 40 KIAS or full down when airspeed is below 40 KIAS. The preferred method
of manually slewing the stabilator up is to use
the cyclic mounted stabilator slew-up switch.
4. LAND AS SOON AS PRACTICABLE.

WARNING
If acceleration is continued or collective is
decreased with the stabilator in a trailing
edge down position, longitudinal control
will be lost. The stabilator shall be slewed
to 0° above 40 KIAS and full-down when
airspeed is less than 40 KIAS. If the stabilator is slewed up to the 0° position and
the AUTO CONTROL switch is pressed
during acceleration, the helicopter may
pitch to a nose down attitude.
Pressing the AUTO CONTROL RESET
button after a failure occurs results in the
automatic mode coming on for one second. If a hardover signal to one actuator
is present, the stabilator could move approximately 4° to 5° in that one second
before another auto mode failure occurs.
Subsequent reset attempts could result in
the stabilator moving to an unsafe position.
If the stabilator AUTO mode repeatedly
disengages during a flight, flight above 70
KIAS is prohibited with the stabilator in
AUTO mode.
If an AUTO Mode Failure Occurs:
NOTE
Use of cyclic mounted stabilator slew-up
switch should be announced to the crew to
minimize cockpit confusion.

9-20

Change 10

If manual control is not possible:
5. STAB POS indicator - Check and fly at or below KIAS LIMITS shown on placard.
6. LAND AS SOON AS PRACTICABLE.
9.31 UNCOMMANDED NOSE DOWN/UP PITCH ATTITUDE CHANGE.
a. An uncommanded nose down/up pitch attitude
change could be the result of a stabilator or other AFCS
malfunction (SAS or FPS). There is a remote possibility
that a stabilator malfunction could occur in the automatic or
manual mode without audio warning or caution light illumination.
b. If an uncommanded nose down pitch attitude change
is detected, the pilot should initially attempt to stop the rate
with aft cyclic. Maintaining or increasing collective position may assist in correcting for a nose down pitch attitude.
If the nose down pitch rate continues, and/or inappropriate
stabilator movement is observed, activate the cyclic
mounted stabilator slew-up switch to adjust the stabilator to
control pitch attitude. Continue to monitor the stabilator
position when the cyclic mounted stabilator slew-up switch
is released to ensure movement stops.
c. Uncommanded nose up pitch attitude changes at airspeeds of 140 KIAS and less should not become severe
even if caused by full up slew of the stabilator and can be
corrected with forward cyclic. If the nose up pitch attitude
is caused by full up stabilator slew at airspeeds above 140
KIAS, full forward cyclic may not arrest the nose up pitch
rate.
d. If an uncommanded nose up pitch attitude change is
detected, the pilot should initially attempt to stop the rate

TM 1-1520-237-10

with forward cyclic. At airspeeds above 140 KIAS, a collective reduction of approximately 3 inches, simultaneously
with forward cyclic will arrest the nose up pitch rate. If
these control corrections are delayed and/or a large nose up
attitude results, a moderate roll to the nearest horizon will
assist in returning the aircraft to level flight. After the nose
returns to the horizon, roll to a level attitude. After coordination with the pilot, the copilot should adjust the stabilator
to 0° at airspeeds above 40 KIAS and full down at airspeeds below 40 KIAS.

4. MAN SLEW switch - Adjust to 0° at airspeeds
above 40 KIAS and full down at airspeeds below 40 KIAS.

5. LAND AS SOON AS PRACTICABLE.
If an uncommanded nose up pitch attitude occurs:
1. Cyclic - Adjust as required.
2. Collective - Reduce as required.

If an uncommanded nose down pitch attitude occurs:
1. Cyclic - Adjust as required.

3. MAN SLEW switch - Adjust to 0° at airspeeds
above 40 KIAS and full down at airspeeds below 40 KIAS.

2. Collective - Maintain or increase.
4. LAND AS SOON AS PRACTICABLE.
3. Cyclic mounted stabilator slew-up switch - Adjust as required to arrest nose down pitch rate.

Change 10

9-21

TM 1-1520-237-10

Section II MISSION EQUIPMENT
9.32 EMERGENCY JETTISONING.
When conditions exist which require the jettisoning of
external loads to ensure continued flight or execution of
emergency procedures, the crew should jettison the load as
follows:
CARGO REL or HOOK EMER REL button
- Press.

9.34.2 PWR MAIN RTR and/or TAIL RTR MONITOR
Light On.
If a PWR monitor light is on with BLADE DEICE
POWER switch ON to stop power from being applied to
blades:
1. Icing conditions - EXIT.
2. BLADE DEICE POWER switch - OFF.

9.33 EMERGENCY RELEASE OF RESCUE HOIST
LOAD.
If the rescue hoist becomes jammed, inoperative, or the
cable is entangled and emergency release is required:

If a PWR monitor light is still on with BLADE DEICE
POWER switch OFF:
3. GENERATORS NO. 1 or NO. 2 switch OFF.

To cut cable from cockpit:
1. CABLE SHEAR switch - FIRE.
To cut cable from hoist operator’s position:
2. CABLE CUT switch - FIRE.
9.34 BLADE DEICE SYSTEM MALFUNCTIONS.
9.34.1 MR DE-ICE FAULT, MR DE-ICE FAIL, or TR
DE-ICE FAIL Caution Light On.
a. If the MR DE-ICE FAULT caution light goes on,
the system will continue to function in a degraded mode.
The pilot must be aware of vibration levels and % TRQ
requirements, which could be a result of ice buildup.

4. APU generator switch - OFF (if in use).

5. LAND AS SOON AS PRACTICABLE.
9.34.3 Ice Rate Meter Fail or Inaccurate. Failure of
the ice rate meter should be indicated by appearance of the
FAIL flag on the meter face. Inaccuracy of the meter will
be indicated by increased torque required and/or increase of
vibration levels due to ice buildup. If failure or inaccuracy
is suspected, with no other indicated failures, the system
can be manually controlled.
1. BLADE DEICE MODE switch - MANUAL
as required.

If vibration levels increase or % TRQ required increases:
2. Higher icing MODE - Select as required.

b. If the MR DE-ICE FAIL caution light goes on, the
main rotor deice will automatically turn off. Tail rotor deice
will remain on.

If ice buildup continues:
3. LAND AS SOON AS PRACTICABLE.

c. If the TR DE-ICE FAIL caution light goes on, tail
rotor deice will automatically turn off. Main rotor deice will
remain on.
1. Icing conditions - Exit.
2. BLADE DEICE POWER switch - OFF, when
out of icing conditions.

If vibrations increase:
3. LAND AS SOON AS POSSIBLE.

9-22

Change 10

9.34.4 Loss of NO. 1 or NO. 2 Generator During
Blade Deice Operation. Loss of one generator during
blade deice operation will result in loss of power to the
system. To restore system operation, the APU must be
started and the APU generator switch ON. The APU GEN
ON advisory light will not go on because one main generator is still operating. The APU generator will supply power
only for blade deice operation.
Pilot not on the controls:
EMER APU START.

TM 1-1520-237-10

9.35 EXTERNAL EXTENDED RANGE FUEL SYSTEM FAILURE TO TRANSFER SYMMETRICALLY.

1. Make all turns shallow (up to standard rate),
and in the direction away from heavy side (particularly when a right tank remains full).

ERFS

2. Avoid abrupt control motions, especially lateral
cyclic.

b. Total failure of one tank to transfer fuel will turn on
the associated tank’s NO FLOW light. Reduced flow from
one tank may not cause a NO FLOW light to go on, but
will change the lateral CG of the helicopter. The pilot will
notice a migration of the lateral cyclic stick position as the
lateral CG offset from neutral increases. For example, a
fully asymmetric outboard 230-gallon tank set (one tank
full, one tank empty), on an otherwise neutrally balanced
helicopter, will result in a level flight lateral stick position
offset of approximately two inches. If asymmetric transfer
is suspected, stop transfer on the selected tank set and initiate transfer on the other tank set, if installed.

4. Select a suitable roll-on landing area, and make
a roll-on landing with touchdown speed in excess of 30 KIAS. To increase control margin,
execute the approach into the wind or with a
front quartering wind from the heavy side and
align the longitudinal axis of the aircraft with
the ground track upon commencing the approach. If a suitable roll-on landing area is not
available, make an approach to a hover into the
wind, or with a front quartering wind from the
heavy side.

If asymmetric fuel transfer is suspected:

3. If possible, shift personnel to the light side of
the helicopter.

9.35A AUXILIARY FUEL MANAGEMENT SYSTEM
FAILURE TO TRANSFER SYMMETRICALLY.
AFMS

1. Stop transfer on tank set.
2. Select other tank set and initiate transfer.
3. LAND AS SOON AS PRACTICABLE.

a. Total failure of a single external extended range fuel
system tank to transfer fuel could be the result of a loose
filler cap, bleed-air regulator/shutoff valve, fuel shutoff
valve, or line blockage failure.
b. Total failure of one tank to transfer fuel will turn on
the associated tank’s NO FLOW light. Reduced flow from
one tank may not cause a NO FLOW light to go on, but
will result in a lateral imbalance between tank pairs whether
AUTO or MAN mode is selected. Occasional monitoring
of the auxiliary fuel management panel (AFMP) fuel quantity displays provides the crew an accurate means of identifying a developing imbalance condition.
When an asymmetric fuel condition (greater than 400
pound difference between external tank pairs) is identified:
1. XFER MODE switch - MAN.

CAUTION

Monitor fuel transfer to remain within
CG limits and avoid asymmetric loading.
2. MAN XFER switch - Select heavy tank LEFT
or RIGHT until imbalance condition is cor-

Change 10

9-23

TM 1-1520-237-10

rected. If the imbalance condition cannot be
corrected:
3. XFER MODE switch - OFF.
4. LAND AS SOON AS PRACTICABLE.

WARNING
With asymmetric fuel loading, lateral control margin will be reduced in the direction opposite the heavy side. The aircraft
has been flown from hover to 138 KIAS,
with lateral CGs equivalent to a fully
asymmetric outboard 230-gallon tank set,
(full right tank, no stores on left side). The
most critical maneuvers are turns toward
the heavy side and approaches with a
crosswind from the lighter side. These
maneuvers are not recommended. The
most adverse condition for lateral controllability is right side heavy, in the 20 to 50
KIAS range. Do not exceed 30° angle of
bank. If controllability is in question, jettison the asymmetric tank set.
Should controlled flight with one heavy external tank
become necessary, proceed as follows:
1. Make all turns shallow (up to standard rate),
and in the direction away from heavy side (particularly when a right tank remains full).
2. Avoid abrupt control motions, especially lateral
cyclic.
3. If possible, shift personnel to the light side of
the helicopter.
4. Select a suitable roll-on landing area, and make
a roll-on landing with touchdown speed in excess of 30 KIAS. To increase control margin,
execute the approach into the wind or with a
front quartering wind from the heavy side and
align the longitudinal axis of the aircraft with
the ground track upon commencing the approach. If a suitable roll-on landing area is not
available, make an approach to a hover into the
wind, or with a front quartering wind from the
heavy side.

9-24

Change 10

9.36 EXTERNAL STORES JETTISON.

At high gross weights and with one engine inoperative,
or in an emergency or performance limited situation, it may
be necessary to jettison a tank set. Circuitry prevents the
release of any individual tank even if a single tank jettison
has been selected at the STORES JETTISON control
panel. The helicopter will remain controllable even if a
single tank fails to release because of a malfunction in the
jettison system. In the case of a four tank configuration, and
depending on the amount of fuel in the tanks, lateral control
may be lost if both tanks on one side fail to release. For this
reason the use of the EMER JETT ALL or JETT ALL
switches is not recommended. Only in circumstances where
failure to do so would result in certain damage to aircraft
and crew, should the use of these switches be considered, at
the discretion of the pilot in command.
If jettisoning of tanks is required:
1. STORES JETTISON switch - Select INBD
BOTH, OUTBD BOTH or ALL as applicable.
2. JETT switch - Actuate.

If primary jettison system does not operate:
3. EMER JETT ALL switch - Actuate.

9.37 FUEL FUMES IN COCKPIT/CABIN WITH EXTERNAL EXTENDED RANGE FUEL SYSTEM
PRESSURIZED. ERFS
If the bleed air check valve(s) is stuck in the open position when the heater is turned on, the resulting bleed air
manifold pressure drops due to the heater bleed air demands. This allows fumes/mist above the tanks to backflow
through the bleed air manifold, through the heater, and into
the cabin. If fuel fumes or mist are noted during external
extended range fuel system operation, perform the following:
If heater is on:
1. HEATER switch - OFF.

If heater is off or fumes persist:
2. PRESS OUTBD and INBD switches - OFF.
3. MODE switch - OFF.

TM 1-1520-237-10

4. FUEL BOOST PUMP CONTROL switches As required.

9.38 FUEL FUMES IN COCKPIT/CABIN WITH EXTERNAL EXTENDED RANGE FUEL SYSTEM
PRESSURIZED. AFMS
If the bleed air check valve(s) is stuck in the open position when the heater is turned on, the resulting bleed air
manifold pressure drops due to the heater bleed air demands. This allows fumes/mist above the tanks to backflow
through the bleed air manifold, through the heater, and into
the cabin. If fuel fumes or mist are noted during external
extended range fuel system operation, perform the following:

3. XFER MODE switch - OFF.
4. FUEL BOOST PUMP CONTROL switches As required.

9.39

VOL

LAUNCHER RACKS JETTISON.

At high gross weights and with one engine inoperative
or in an emergency, it may be necessary to jettison the
volcano launcher racks. Both lower launcher racks must
separate from helicopter before upper racks are activated. If
one lower rack remains, upper racks will not jettison.
If jettisoning of launcher rack is required:
1. JETTISON switch - JETTISON.

If heater is on:
1. HEATER switch - OFF.

If heater is off or fumes persist:

If jettison procedure above fails, do the following immediately:
2. EMER JETTISON switch - JETTISON.

2. PRESS switch - OFF.

Change 4

9-25/(9-26 Blank)

TM 1-1520-237-10

APPENDIX

REFERENCES

AR 70-50

Designating and Naming Military Aircraft, Rockets and Guided Missiles

AR 95-1

Army Aviation General Provisions and Flight Regulations

AR 95-3

General Provisioning, Training, Standardization, and Resource Management

AR 385-40

Accident Reporting and Records

DA PAM 40-501

Noise and Conservation of Hearing

DA PAM 738-751

Functional Users Manual for the Army Maintenance Management System Aviation
(TAMMS-A)

DOD FLIP

Flight Information Publication

FAR Part 91

Federal Air Regulation, General Operating and Flight Rules

FM 1-202

Environmental Flight

FM 1-203

Fundamentals of Flight

FM 1-230

Meteorology for Army Aviators

FM 1-240

Instrument Flying and Navigation for Army Aviators

FM 10-68-1

Aircraft Refueling

FM 55-450-1

Army Helicopter External Load Operations

FM 55-450-2

Army Helicopter Internal Load Operations

TB 55-9150-200-24

Engine and Transmission Oils, Fuels, and Additives for Army Aircraft

TC 1-204

Night Flight Techniques and Procedures

TM 1-1520-237-CL

Operator’s and Crewmember’s Checklist, Army Model UH-60A, UH-60L and EH-60A
Helicopters

TM 1-1520-250-23

General Tie-down and Mooring Technical Manual Aviation Unit and Intermediate
Maintenance. All Series Army Models AH-64, UH-60, CH-47, UH-1, AH-1, OH-58
Helicopters.

TM 9-1005-224-10

Operator’s Manual M60 Machinegun

TM 9-1300-206

Ammunition and Explosives

TM 11-5810-262-10

Operator’s Manual for Communication Security Equipment TSEC/KY 58

TM 11-5841-281-12

Operator’s Manual for Doppler Navigation Set

TM 11-5841-283-12

Operator’s Manual for AN/APR 39(V) Radar Signal Detecting Set

TM 11-5865-200-12

AN/ALQ 144 Countermeasure Set

TM 11-5895-1199-12

Operator’s and Organization Maintenance for MARK-12 IFF System AN/APX-100,
AN/APX-72

Change 6

A-1

TM 1-1520-237-10

APPENDIX

(Cont)

TM 11-5895-1037-12 & P

Transponder Set AN/APX-100(V)

TM 55-1500-342-23

Army Aviation Maintenance Engineering Manual: Weight and Balance

TM 750-244-1-5

Procedures for the Destruction of Aircraft and Associated Equipment to Prevent Enemy
Use

A-2

Change 1

TM 1-1520-237-10

APPENDIX

ABBREVIATIONS AND TERMS

- auxiliary fuel management
panel

ECU

- electrical control unit

EMB

- expanded memory board

AFMS

- auxiliary fuel management
system

ENG

- engine

- anti-jam

EOT

- element-on-time

ALT

- altitude

ERFS

- extended range fuel system

ANT

- antenna

ESU

- electronic sequence unit

APU

- auxiliary power unit

ESSS

- external stores support system

ATF

- aircraft torque factor

ETF

- engine torque factor

- butt line

ETL

- effective translational lift

°C

- degree Celsius

°F

- degree Fahrenheit

CBIT

- continuous built in test

FAT

- free-air temperature

CCU

- converter control unit

FPM

- feet-per-minute

CDU

- central display unit

- feet

- center of gravity

GPS

- global positioning system

- center line

- gross weight

CRT

- cathode ray tube

HAT

- height above terrain

- continuous wave

HMU

- hydromechanical unit
(fuel control)

DCU

- dispenser control unit

- have quick

DEC

- digital electronic control (for
engine)

- hour

DEG

- degree

HSI

- horizontal situation indicator

- direction find

HSP

- hot start preventor

DRVS

- Doppler radar velocity sensor

HSS

- horizontal stores support

- display unit

HUD

- heads up display

ECM

- electronic counter measure

IAS

- indicated airspeed

EGR

- embedded GPS receiver

IAW

- in accordance with

- inboard

IBIT

- initiated built in test

AFMP

- change in flat plate drag area

TRQ

- change in torque

Change 7

B-1

TM 1-1520-237-10

APPENDIX

(Cont)

IGE

- in ground effect

NVG

- night vision goggles

IINS

- integrated inertial navigation
system

- degree

% RPM R

- rotor rpm, percent

- inch

% RPM 1 or 2

IN HG

- inch of mercury

- No. 1 or No. 2 engine Np %
rpm

IPS

- inlet particle separator
- inches per second

OAT

- outside air temperature

- outboard

ODV

- overspeed and drain valve

OEI

- one engine inoperative

OGE

- out of ground effect

- pressure altitude

PAS

- power available spindle

PBIT

- power up built in test

PDU

- pilot display unit

POS

- position

POU

- pressurizing and overspeed
unit

- infrared

IRCM

- infrared countermeasures

KCAS

- knots calibrated airspeed

KIAS

- knots indicated airspeed

- knot

KTAS

- knots true airspeed

- pound(s)

lb/gal

- pounds-per-gallon

lb/hr

- pounds-per-hour

LCD

- liquid crystal display

LDI

- leak detection isolation

PRESS

- pressure

LDS

- load-demand spindle

PPM

- pounds-per-minute

LIM

- limit

PSCU

- power supply calibration unit

LLL

- low light level

PSI

- pounds per square inch

LRU

- line replaceable unit

PSID

LWC

- liquid water content

- pounds per square inch
differential

MAX

- maximum

PSIG

- pounds per square inch
gauge

MGRS

- military grid reference system

R/C

- rate of climb

MIN

- minimum

R/D

- rate of descent

min

- minutes

RDW

- ram dump waypoints

Ng SPEED 1 or 2

- No. 1 or No. 2 engine
compressor speed % rpm

RPM

- revolutions-per-minute

- nautical miles

RTA

- Receiver Transmitter Antenna

- power turbine speed

SDC

- signal data converter

B-2

Change 8

TM 1-1520-237-10

APPENDIX

SEL

- select

- sea level

SLAB

- sealed lead acid battery

SPEC

- specification

STA

- station

STD TEMP

- 15°C at sea level

SQ FT

- square feet

TAS

- true airspeed

TGT

- turbine gas temperature

TOD

- time of day

% TRQ

- torque, percent

TRQ

- torque

UTM

- universal transverse mercator

VDC

- volts direct current

(Cont)

- maximum level flight speed
using torque available - 30
minutes

VIDS

- vertical instrument display
systems

Vne

- velocity never exceed
(airspeed limitation)

VSI

- vertical situation indicator

VSP

- vertical support pylon

- water line

WOD

- word of day

WOW

- weight-on-wheels

XMSN

- transmission

Change 8

B-3/(B-4 Blank)

TM 1-1520-237-10

APPENDIX

KY-100.
C.1 KY-100 CONFIGURATION SETUP.
CONTROL/
DISPLAY
d and g

INIT

FUNCTION
Used to scroll through the available
menus or available options. g
scrolls in the opposite direction of
the d. Pressing d and g
simultaneously returns the display
to the previous set of menus (one
level up from the current menu on
the display).
Used to bring up the set of menus
or options that are one level below
the current menu on the display.

(10) Repeat steps (6) through (9) until all of the
menu items shown in Table C-1 have been set
as indicated.
(11) After all menu items from Table C-1 have been
correctly set, d and g - Press simultaneously,
and AUd-dATA will be displayed.
(12) d and g - Press simultaneously, and INFC will
be displayed. Proceed to step b. to set Radio
Interface settings.
b. Radio Interface.
(1) PRESET switch - MAN.
(2) MODE switch - OFL.

C.1.1 KY-100 Audio/Data and Radio Interface Settings Procedure.

(3) d - Repeatedly press until INFC is displayed.
(4) INIT - Press, and AUd-dATA will be displayed.

a. Audio Data Interface.
(5) d or g - Press until RAdIO is displayed.
(1) PRESET switch - MAN.
(2) MODE switch - OFL.

(6) INIT - Press, and NRW-bANd will be displayed.

(3) d - Repeatedly press until INFC is displayed.

(7) INIT - Press.

(4) INIT - Press, and AUd-dATA will be displayed (if necessary use d or g to display
AUd-dATA).

(8) d or g - Press until SET dEF is displayed.
(9) INIT - Press, and a flashing SET dEF will be
displayed.

(5) INIT - Press.
(6) Select the item that needs to be changed (Menu
Item column of Table C-1) by repeatedly pressing d until the item is displayed.

(10) INIT - Press. SET dEF stops flashing and a
pass tone is heard, indicating nrw-band radio
default settings are stored in memory.
(11) d and g - Press simultaneously.

(7) INIT - Press, and the default setting will be
displayed.
(8) If the setting does not agree with the setting
shown under the Setting column of Table C-1,
d - Press until the proper setting is displayed.
(9) d and g - Press simultaneously.

(12) Select the item that needs to be changed (Menu
Item column of Table C-2) by repeatedly pressing d or g until the item is displayed.
(13) INIT - Press, and the default setting or a SubMenu (if one exists) will be displayed. Unless
Table C-2 indicates a SubMenu, go to step (15).

Change 8

C-1

TM 1-1520-237-10

APPENDIX

(Cont)

Table C-1.
MENU ITEM

SETTING

DEFAULT

GUARd

GRd OFF

YES

MIC

MIC UNbAL

YES

bALANCE

RX UNbAL*

YES

IMPEd

150 OHMS

dAT SENS

MARK +

YES

RX COUP

RX AC

YES

TX COUP

TX AC

YES

TX CLK

J2-V

* For MH Series aircraft MIC should be set
to MIC BAL.
(14) In cases with a SubMenu, select the appropriate
SubMenu by pressing d until the SubMenu indicated in Table C-2 is displayed. Then press
INIT, and the default setting will be displayed.
d - Press until proper setting is displayed. d
and g - Press simultaneously. Repeat this step
for all SubMenus within the same Menu Item.
When all SubMenu settings within the same
Menu Item have been set, d and g - Press simultaneously.

(15) If the setting does not agree with the setting
shown under the Setting column of Table C-2,
d or g - Press until the proper setting is displayed.
(16) d and g - Press simultaneously.
(17) Repeat steps (12) through (16) until all of the
menu items shown in Table C-2 have been set
as indicated.

Table C-2.

C-2

MENU ITEM

SUBMENU

SETTING

DEFAULT

TX CLKS

---------

EXT CLK

TRN SEQ

---------

YES

TX dELAY

---------

135 ms

YES

PREAM

---------

ENH

dAT SENS

---------

MARK -

YES

CTS

Bd/BdL

188

CTS

HF/PT

188

CTS

LOS

188

MILSTAR

---------

OFF

YES

TX LVL

---------

YES

IMPEd

---------

150 OHMS

Change 8

TM 1-1520-237-10

APPENDIX

(Cont)

Table C-2. (Cont)
MENU ITEM

SUBMENU

SETTING

DEFAULT

RTS/PTT

Bd/BdL

RTS

RTS/PTT

PTT

RTS/PTT

LOS

PTT

RTS/PTT

PTT

C.1.2 KY-100 Preset Configuration Procedure.

i. d or g - Press until BD is displayed. Press
INIT to select this setting.

a. MODE switch - OFL.
b. PRESET switch - Rotate to preset to be modified (1 - 6).
c. d or g - Press until PRESET is displayed.
d. INIT - Press. If NARROWBAND is not displayed, INIT - Press. The preset operating
mode will flash.

j. Press d or g until RATE SELECT is displayed. Press INIT to select this setting.
k. d or g - Press until RATE 24 is displayed.
INIT - Press to select this setting.
l. d or g - Press until KEY SELECT is displayed. INIT - Press to select this setting.

e. d or g - Press until NRW-BAND is displayed.

m. d or g - Press until TEK 1 is displayed. INIT
- Press to select this setting.

f. INIT - Press, NARROWBAND is now selected.

n. d or g - Press until MODE SELECT is displayed. INIT - Press to select this setting.

g. d or g - Press until MODEM SELECT is
displayed.

o. dor g - Press until HF VC VT is displayed.
INIT - Press to select this setting.

h. INIT - Press.

Change 8

C-3/(C-4 Blank)

TM 1-1520-237-10

INDEX
Subject

Paragraph
Figure, Table
Number

#1 and #2

#1 and #2 CONV Caution Lights On ...............................................................................................................

9.26.3

#1 or #2 FUEL FLTR BYPASS Caution Light On .........................................................................................

9.25.1

#1 and #2 FUEL FLTR BYPASS Caution Lights On .....................................................................................

9.25.2

#1 or #2 HYD PUMP Caution Light On and BACK-UP PUMP ON Advisory Light Off............................

9.27.4

#1 HYD PUMP Caution Light On....................................................................................................................

9.27.1

#1 or #2 PRI SERVO PRESS Caution Light On .............................................................................................

9.27.5

#1 RSVR LOW and #1 HYD PUMP Caution Lights On With BACK-UP PUMP
ON Advisory Light On ......................................................................................................................................

9.27.6

#1 FUEL LOW and #2 FUEL LOW Caution Lights On.................................................................................

9.25.3

#1 or #2 FUEL PRESS Caution Light On........................................................................................................

9.25.4

#1 and #2 HYD PUMP Caution Lights On......................................................................................................

9.27.3

#1 and #2 Generator Failure (#1 and #2 CONV and AC ESS BUS OFF
Caution Lights On).............................................................................................................................................

9.26.1

#1 or #2 GEN Caution Light On.......................................................................................................................

9.26.2

#2 RSVR LOW Caution Light On....................................................................................................................

9.27.8

#2 RSVR LOW and #2 HYD PUMP Caution Lights On With BACK-UP PUMP
ON Advisory Light On ......................................................................................................................................

9.27.7

#2 HYD PUMP Caution Light On....................................................................................................................

9.27.2

#1 TAIL RTR SERVO Caution Light On and BACK-UP PUMP ON Advisory light Off or #2 TAIL RTR
SERVO ON Advisory Light Off .......................................................................................................................

9.22.6

A
AC Power Supply System .................................................................................................................................

2.65
2.65.2

AC Secondary Bus

.......................................................................................................................................

Accept/Reject Page

.......................................................................................................................................

3-27

Accessory Module..............................................................................................................................................

2.45.2

Accumulator Recharge .......................................................................................................................................

2.69

After Emergency Action ....................................................................................................................................

9.4

After Landing Check..........................................................................................................................................

8.33

After Takeoff ......................................................................................................................................................

8.30

................................................................................................................................

2.61

Air Conditioner System

Air Conditioning System Power Priority

.....................................................................................................

Change 1

2-2

INDEX-1

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Aircraft Configuration Drag Changes For Use With Clean Cruise Charts......................................................

7.21
7A.22

Aircraft Configuration Drag Changes For Use With High Drag Cruise Charts..............................................

7.22
7A.23

Aircraft Torque Factor (ATF)............................................................................................................................

7-2

7A-2

Airspeed Charts ..................................................................................................................................................

7.28
7A.29

Airspeed Correction Aircraft With Wedge Mounted Pitot-Static Probes ........................................................

7-36

Airspeed Correction Aircraft Without Wedge Mounted Pitot-Static Probes ...................................................

7-35

Airspeed Correction Chart .................................................................................................................................

7A.38

Airspeed Correction Chart - High Drag............................................................................................................

7-37

7A-39

Airspeed Correction Charts................................................................................................................................

7.28.1
7A.29.1

Airspeed for Onset of Blade Stall .....................................................................................................................

5-9

Airspeed Indicator ..............................................................................................................................................

2.75

Airspeed Limitations Following Failure of the Automatic Stabilator Control System ...................................

5.21

Airspeed Operating Limits.................................................................................................................................

5-6

...........................................................................................................................

5-7

Airspeed Operating Limits

5.19
Airspeed Operating Limits (Volcano) ...............................................................................................................

5-8
7A.28

Airspeed System Description.............................................................................................................................

7.27
7A.28

Airspeed System Dynamic Characteristics........................................................................................................

7.28.2
7A.29.2

Altimeter Encoder AAU-32A ............................................................................................................................

2-22
2.76

Altimeter Set AN/APN-209 ...............................................................................................................................

Altimeter Set AN/APN-209(V) .........................................................................................................................
AN/APR-39(V)2 Control Panel .........................................................................................................................

INDEX-2

Change 1

3-29
3.26

4-5

TM 1-1520-237-10

INDEX (Cont)
Subject

AN/ASN-132(V)

Paragraph
Figure, Table
Number

...........................................................................................................................................

Antenna Arrangement ........................................................................................................................................

3.16
F

3-1

Anticollision Lights............................................................................................................................................

2.71.3

Appendix A, References ....................................................................................................................................

1.4

Appendix B, Abbreviations and Terms.............................................................................................................

1.5

Approved Fuels ..................................................................................................................................................

2-5

APU ...................................................................................................................................................................

2.68

APU Compartment Fire .....................................................................................................................................

9.24.2

APU Controls ....................................................................................................................................................

2.68.1

APU Fuel Control System (Helicopters equipped with GTC-P36-150 APU).................................................

2.68.3

APU Fuel Control System (Helicopters equipped with T-62T-40-1 APU).....................................................

2.68.2

APU Fuel Supply System ..................................................................................................................................

2.68.4

APU Inlet Particle Separator (IPS) Kit. (Helicopters with IPS kit installed)..................................................

4.26

APU Oil System Servicing ................................................................................................................................

2.87

APU OIL TEMP HI Caution Light On.............................................................................................................

9.23.3

APU Operating Limitations ...............................................................................................................................

5.31

APU Source Engine Start ..................................................................................................................................

2.28.2

Arm .....................................................................................................................................................................

6.6.2

Armament Loading Data Moments ...................................................................................................................

6-5
6.13

Armament Subsystem.........................................................................................................................................
Arming Volcano System Canisters....................................................................................................................

4.14
F

Army Aviation Safety Program .........................................................................................................................
ASE Status Panel

..........................................................................................................................................

4-21
1.7

4-10
4.13

Assault Mission Profile (4 - 230 Gallon Tanks)...............................................................................................

Assault Mission Profile (2 - 230 Gallon Tanks)...............................................................................................

7-39

7A-41

7-40

7A-42

Attitude Indicating System.................................................................................................................................
Authorized Ammunition.....................................................................................................................................
Automatic Flight Control System (AFCS)........................................................................................................

Change 1

2.73
T

4-2
2.37

INDEX-3

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Automatic Flight Control System (AFCS) Switch Panel .................................................................................

2-15

Autorotative Glide Distance Chart ...................................................................................................................

9-4

Autorotative Glide Distance Chart-High Drag..................................................................................................

9-5

Auxiliary AC Power System .............................................................................................................................
Auxiliary Cabin Heater Control Panel ..............................................................................................................

2.66
F

Auxiliary Electrical Cabin Heater (On helicopters equipped with auxiliary cabin heater kit) .......................
Auxiliary Fuel Management Control Panel.......................................................................................................

4-26
4.20

4-27
4.20.3

Auxiliary Fuel Management Control Panel

AFMS

.....................................................................................

4-27.1
4.22A.2

Auxiliary Fuel Management Control Panel Test ..............................................................................................
Auxiliary Fuel Management System Fault Messages
Auxiliary Heater System.

.....................................................................

AFMS

4.20.5
T

4-4

.........................................................................................................................

2.62

Auxiliary Power Unit (APU) System................................................................................................................

2.67

Auxiliary Power Unit (APU) (Typical).............................................................................................................

2-21

Average Arm ......................................................................................................................................................

6.6.4

Aviation Life Support Equipment .....................................................................................................................

8.3

Avionics Equipment Configuration ...................................................................................................................

3.2

Avionics Power Supply......................................................................................................................................

3.3

B
Backup Hydraulic Pump Hot Weather Limitations ..........................................................................................

5.30

Backup Hydraulic System..................................................................................................................................

2.40.3

Balance Definitions ............................................................................................................................................

6.6

Bank Angle Limitation.......................................................................................................................................

5.23.3.4

Basic Moment.....................................................................................................................................................

6.6.5

BATT LOW CHARGE Caution Light On .......................................................................................................

9.26.5

Battery.................................................................................................................................................................

2.64.2

Battery Charger/Analyzer...................................................................................................................................

2.64.5

BATTERY FAULT Caution Light On .............................................................................................................

9.26.4

..........................................................................................

4.5

Before Exterior Check ......................................................................................................................................

8.11

Before Landing...................................................................................................................................................

8.31

Bearing, Distance, Heading Indicator (BDHI)

INDEX-4

Change 5

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Before Leaving Helicopter.................................................................................................................................

8.35

Before Starting Engines .....................................................................................................................................

8.21

Before Takeoff....................................................................................................................................................

8.28

Before Taxi.........................................................................................................................................................

8.25

Blackout Curtains ...............................................................................................................................................

2.55

Blade Deice System Control Panel ...................................................................................................................

2.54.2

Blade Deice System Malfunctions.....................................................................................................................

9.34

Blade Deice System Operation..........................................................................................................................

2.54.1

Blade Deice Test ................................................................................................................................................

2.54.3

Blade Deice Test Panel......................................................................................................................................

2.54.4

Boost Servo ........................................................................................................................................................

2.36.4

BOOST SERVO OFF Caution Light On..........................................................................................................
Built In Test (BIT) Chip Detectors ...................................................................................................................

2.46.3

C
Cabin Ceiling Tiedown Fittings.........................................................................................................................
Cabin Dimensions ..............................................................................................................................................

5.16
F

6-9
6.15

Cabin Dome Lights ............................................................................................................................................

2.70.7

Cabin Doors........................................................................................................................................................

6.16

Cabin Interior .....................................................................................................................................................

2-5

..................................................................................................

2-6

Cabin Mission Equipment Arrangement

Cabin Top (AREA 3).........................................................................................................................................

8.15

Cargo Center of Gravity Planning.....................................................................................................................

6.20.2

Cargo Hook Moments........................................................................................................................................

6-7

Cargo Hook System ...........................................................................................................................................

4-23
4.17

Cargo Hook Weight Limitation .........................................................................................................................
Cargo Tiedown Arrangement.............................................................................................................................

5.17
F

Caution/Advisory BRT/DIM - TEST Switch....................................................................................................
Caution/Advisory and Warning Light Lighting Parameters .............................................................................

2.81.2
T

Caution/Advisory Light System.........................................................................................................................
Caution/Advisory Panel

.............................................................................................................................

Change 5

6-11

2-3
2.81.1

2-24

INDEX-5

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

CDU and PDU Digital Control .........................................................................................................................
CDU Controls and Indicators

....................................................................................................................

2.10.4.3
F

3-19

Center of Gravity (CG)......................................................................................................................................

6.6.6

Center of Gravity Limitations............................................................................................................................

5.13

Center of Gravity Limits Chart .........................................................................................................................

6-13
6.21

Center of Gravity Limits Chart

.................................................................................................................

6-14

Central Display Unit (CDU)..............................................................................................................................

2.10.4

CG Limits ...........................................................................................................................................................

6.6.7

Chaff and Flare Dispenser M130.......................................................................................................................

4.3

Chaff Dispenser M130 .......................................................................................................................................

4.3.1

Chapter 7 Index..................................................................................................................................................

7.2

Chapter 7A Index ...............................................................................................................................................

7A.2

Charging (Cocking) Machinegun M60D...........................................................................................................

4-18

Checklist .............................................................................................................................................................

8.9

CHIP INPUT MDL LH or RH Caution Light On ...........................................................................................

9.22.8

CHIP MAIN MDL SUMP, CHIP ACCESS MDL LH or RH, CHIP TAIL XMSN or CHIP INT
XMSN/TAIL XMSN OIL TEMP or INT XMSN OIL TEMP Caution Light On..........................................

9.22.9

CIS Modes of Operation....................................................................................................................................

Class....................................................................................................................................................................
Climb/Descent ....................................................................................................................................................

3-32
6.2

7-32

7A-35

Climb/Descent Chart ..........................................................................................................................................

7.23
7A.24

Climb/Descent - High Drag ...............................................................................................................................

7-33

7A-36

Clock...................................................................................................................................................................
Cockpit Diagram ................................................................................................................................................

2.80
F

2-4

Cockpit Doors.....................................................................................................................................................

2.11.1

Cockpit Equipment Checks................................................................................................................................

8.22

Cockpit Floodlights ............................................................................................................................................

2.70.2

Cockpit-Left Side (AREA 2) .............................................................................................................................

8.14

INDEX-6

Change 5

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Cockpit-Right Side (AREA 8)...........................................................................................................................

8.20

Cold Weather Control Exercise .........................................................................................................................

8.41.2

Cold Weather Operation ....................................................................................................................................

8.41

Cold Weather Preflight Check ...........................................................................................................................

8.41.1

Collective/Airspeed to Yaw (Electronic Coupling) ..........................................................................................

2.35.4

Collective and Cyclic Grips...............................................................................................................................

2-14

Collective Boost Servo Hardover/Power Piston Failure...................................................................................

9.27.9

Collective Bounce/Pilot Induced Oscillation.....................................................................................................

8.39.3

Collective Pitch Control Stick ...........................................................................................................................

2.35.2

Communication/Navigation Equipment.............................................................................................................

3-1

Compartment Diagram

...............................................................................................................................

2.7.2

Compartment Diagram

...............................................................................................................................

2.7.1

Compass Control Panel C-8021/ASN-75 ..........................................................................................................

3-29

Configuration Drag Change ...............................................................................................................................

7-1

7A-1

Configurations

............................................................................................................................................

5.34

Control Display Unit AN/ARC-220 ..................................................................................................................

3-12.1

Control Panel KY-100(V)..................................................................................................................................

3-12.2

Control Panel RT-1296/APX-100(V) ................................................................................................................

3-33

Converters...........................................................................................................................................................

2.64.1

...................................................................................................

4.10

.....................................................................................................

4.11

Crash-Actuated System ......................................................................................................................................

2.14.6

Crash Axe ...........................................................................................................................................................

2.15

Crew Briefing .....................................................................................................................................................

8.5

Crew Chief/Gunner Seats...................................................................................................................................

2.12.3

Crew Chief/Gunner Windows............................................................................................................................

2.11.3

.......................................................................................................................

4.2

Crew Duties/Responsibilities .............................................................................................................................

8.4

Crew Seats..........................................................................................................................................................

2.12

Countermeasures Set AN/ALQ-156(V)2
Countermeasures Set AN/ALQ-162(V)

Crew Call Switch/Indicator

Crewmember’s Cargo Hook Control Pendant...................................................................................................

4-24
4.15.3

Change 5

INDEX-7

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Crossbleed Engine Start System........................................................................................................................
Cruise-Pressure Altitude Sea Level ...................................................................................................................

2.28.3
F

7-7

7A-9

Cruise - Pressure Altitude 2,000 Feet ..............................................................................................................

7-9

Cruise - Pressure Altitude 2,000 Feet ..............................................................................................................

7A-11

Cruise - Pressure Altitude 4,000 Feet ...............................................................................................................

7-11

7A-13

7-13

7A-15

7-15

7A-17

7-17

7A-19

7-19

7A-21

7-21

7A-23

7-23

7A-25

7-25

7A-27

7-27

7A-29

7-8

7A-10

7-10

7A-12

7-12

7A-14

7-14

7A-16

Cruise - Pressure Altitude 6,000 Feet ...............................................................................................................

Cruise - Pressure Altitude 8,000 Feet ...............................................................................................................

Cruise - Pressure Altitude 10,000 Feet .............................................................................................................

Cruise - Pressure Altitude 12,000 Feet .............................................................................................................

Cruise - Pressure Altitude 14,000 Feet .............................................................................................................

Cruise - Pressure Altitude 16,000 Feet .............................................................................................................

Cruise - Pressure Altitude 18,000 Feet .............................................................................................................

Cruise - Pressure Altitude 20,000 Feet .............................................................................................................

Cruise High Drag - Pressure Altitude Sea Level..............................................................................................

Cruise High Drag - Pressure Altitude 2,000 Feet.............................................................................................

Cruise High Drag - Pressure Altitude 4,000 Feet.............................................................................................

Cruise High Drag - Pressure Altitude 6,000 Feet.............................................................................................

INDEX-8

Change 5

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Cruise High Drag - Pressure Altitude 8,000 Feet.............................................................................................

Cruise High Drag - Pressure Altitude 10,000 Feet...........................................................................................

Cruise High Drag - Pressure Altitude 12,000 Feet...........................................................................................

Change 5

7-16

7A-18

7-18

7A-20

7-20

INDEX-8.1/(INDEX-8.2 Blank)

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Cruise High Drag - Pressure Altitude 12,000 Feet...........................................................................................

7A-22

Cruise High Drag - Pressure Altitude 14,000 Feet...........................................................................................

7-22

7A-24

7-24

7A-26

7-26

7A-28

7-28

7A-30

Cruise High Drag - Pressure Altitude 16,000 Feet...........................................................................................

Cruise High Drag - Pressure Altitude 18,000 Feet...........................................................................................

Cruise High Drag - Pressure Altitude 20,000 Feet...........................................................................................

Cryptographic Computer Kit-1C .......................................................................................................................

3.28

Cyclic-Mounted Stabilator Slew Up Switch .....................................................................................................

2.38.3

Cyclic Stick ........................................................................................................................................................

2.35.1

D
Dashed Line Data...............................................................................................................................................

7.5.1

Data Basis...........................................................................................................................................................

7.5.2
7A.5.1

Data Compartments............................................................................................................................................
DC and AC Circuit Breaker Panels (Typical) ..................................................................................................

2.58
F

2-20
2.64.6

.........................................................................................................................................

2.64.3

DC Power Supply System .................................................................................................................................

2.64

DD Form 365-3 (Chart C) Weight and Balance Records ................................................................................

6.7

DD Form 365-4 (Form F)..................................................................................................................................

6.9

Decreasing % RPM R........................................................................................................................................

9.14

DC Monitor Bus

Definition of Course Terms ...............................................................................................................................

3-18.2

Definition of Emergency Terms ........................................................................................................................

9.3

Desert and Hot Weather Operation ...................................................................................................................

8.42

Destruction of Army Materiel to Prevent Enemy Use .....................................................................................

1.8

..................................................................................................................

2.13.2

..........................................................................................................

2.29.2

Dimensions .........................................................................................................................................................

2.5

Direction Finder Set AN/ARN-89 (LF/ADF) ...................................................................................................

3.13

DF and ECM Operator’s Seats

Digital Electronic Control (DEC)

701C

Change 1

INDEX-9

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Direction Finder Set AN/ARN-149 (LF/ADF) .................................................................................................

3.14

Discriminator Off Mode Display .......................................................................................................................

4-2

Discriminator On Mode Display .......................................................................................................................

4-3

Ditching-Power Off ............................................................................................................................................

9.28.3

Ditching-Power On.............................................................................................................................................

9.28.2

Door Locks .........................................................................................................................................................

2.11.4

Doors and Windows...........................................................................................................................................

2.11

............................................................................

3.17A

Doppler/GPS Navigation Set (DGNS) AN/ASN-128B

Doppler/GPS Navigation Set AN/ASN-128B ...................................................................................................

3-18.1

Doppler World UTM Spheroids ........................................................................................................................

3-22

Doppler Lamp Test Mode Display....................................................................................................................

3-18

......................................................................................................

3-17

Doppler Navigation Set AN/ASN-128

3.17
Downwind Hovering ..........................................................................................................................................

5.23.2

Dual-Engine Failure ...........................................................................................................................................

9.13

Dual-Engine Failure-General .............................................................................................................................

9.12

Dual-Engine Fuel Flow......................................................................................................................................

7.26
7A.27

Dual-Engine Torque Limitations at Airspeeds Above 80 KIAS

..........................................................

5-4

Dual-Engine Torque Limit.................................................................................................................................

7A-5

701C

Dual-Engine Torque Limits ...............................................................................................................................

7A.14

E
................................................................................................................................

4.9

Effects of Blade Erosion Kit..............................................................................................................................

7.15

ECM Antenna Switch

7A.16
.......................................................................................................................................................

2.4

EH-60A Helicopters Without Mission Equipment ..........................................................................................

6.14

..............................................................................................................................................

8.2

EH-60A

EH-60A Data

Electrical Control Unit (ECU)

..................................................................................................................

2.29.1

Electrical Fire In Flight......................................................................................................................................

9.23.5

Electrical Power Systems...................................................................................................................................

2.63

700

Electrical System................................................................................................................................................

INDEX-10

Change 1

2-19

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

9.26
Electronic Navigation Instrument Display System ...........................................................................................

3.25

Emergency Equipment (Portable)......................................................................................................................

9.6

Emergency Exits.................................................................................................................................................

9.5

Emergency Exits and Emergency Equipment Diagram....................................................................................

9-1

Emergency Jettisoning .......................................................................................................................................

9.32

Emergency Landing In Wooded Areas. Power Off ..........................................................................................

9.28.1

Emergency Release of Rescue Hoist Load .......................................................................................................

9.33

Engine .................................................................................................................................................................

2.17

......................................................................................................................................

2.19.1

....................................................................................................................................

2.19.2

Engine and Engine Inlet Anti-Ice Limitations ..................................................................................................

5.29

Engine Anti-Icing ...............................................................................................................................................

2.26.1

Engine Anti-Icing Systems ................................................................................................................................

2.26

Engine Bleed Air ..............................................................................................................................................

7.12

Engine Alternator

700

Engine Alternator

701C

7A.12
Engine Bleed-Air System...................................................................................................................................

2.25

Engine Chip Detector.........................................................................................................................................

2.27.4

Engine Compressor Stall....................................................................................................................................

9.18

Engine Control Quadrant ...................................................................................................................................

2-13
2.29.3

Engine Control System ......................................................................................................................................

2.29

Engine Driven Boost Pump ...............................................................................................................................

2.18.1

Engine Emergency Oil System..........................................................................................................................

2.27.1

Engine Fire In Flight..........................................................................................................................................

9.23.4

Engine Fuel Prime System.................................................................................................................................

2.33

Engine Fuel Supply System...............................................................................................................................

2.18

Engine Fuel System Components......................................................................................................................

2.18.4

Engine Fuel System Selector Control ...............................................................................................................

2.32.2

Engine/Fuselage Fire On Ground ......................................................................................................................

9.23.1

Engine High-Speed Shaft Failure ......................................................................................................................

9.20

Engine Inlet Anti-Icing ......................................................................................................................................

2.26.2

Change 1

INDEX-11

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Engine Instruments.............................................................................................................................................

2.31

Engine Limitations .............................................................................................................................................

5.8

Engine Malfunction-Partial or Complete Power Loss ......................................................................................

9.7

ENGINE OIL FILTER BYPASS Caution Light On, ENGINE CHIP Caution Light On,
ENG OIL PRESS HIGH/LOW, ENGINE OIL TEMP HIGH, ENGINE OIL TEMP Caution
Light On, ENGINE OIL PRESS Caution Light On.........................................................................................

9.19

Engine Oil Pressure Indicator............................................................................................................................

2.31.2

Engine Oil System .............................................................................................................................................

2.27

Engine Oil System Characteristics ....................................................................................................................

8.41.4

Engine Oil System Servicing.............................................................................................................................

2.86

Engine Oil Temperature Indicator .....................................................................................................................

2.31.1

Engine Operation................................................................................................................................................

8.41.3

Engine Overspeed Check Limitations ...............................................................................................................

5.11

Engine Power Limitations

700

.......................................................................................................................

5.8.1

Engine Power Limitations

701C

......................................................................................................................

5.8.2

Engine Power Turbine/Rotor Speed Indicator ..................................................................................................

2.31.5

Engine Restart During Flight.............................................................................................................................

9.11

Engine % RPM Limitations...............................................................................................................................

5.8.3

Engine Runup .....................................................................................................................................................

8.24

Engine Speed Control System ...........................................................................................................................

2.29.5

Engine Start Envelope........................................................................................................................................

5-5

Engine Start Limits ............................................................................................................................................

5.10

Engine Starter Limits .........................................................................................................................................

5.8.4

Engine Start System ...........................................................................................................................................

2.28

Engine T700 ......................................................................................................................................................

2-11

Equipment Loading and Unloading...................................................................................................................

6.20

Equipment Stowage Compartments...................................................................................................................

6.19

ESSS Fuel System Degraded Operation Chart

........................................................................................

4-3

Exceeding Operational Limits ...........................................................................................................................

5.3

Explanation of Change Symbols .......................................................................................................................

1.10

Exterior Check....................................................................................................................................................

8.12

Exterior Check Diagram ....................................................................................................................................

INDEX-12

Change 1

8-1

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Exterior Lights....................................................................................................................................................

2.71

External AC Power System ...............................................................................................................................

2.66.2

External Air Source/Electrical Requirements....................................................................................................

2.85

.................................................................................

4.22A

External Extended Range Fuel Flow Verification in Manual Mode................................................................

4.20.7.3

External Extended Range Fuel Quantity Indicating System ............................................................................

4.20.4

External Auxiliary Fuel Management System

AFMS

External Extended Range Fuel System Degraded Operation Chart

ERFS

...............................................

ERFS

4-3

......

9.35

.....................................................................................

4.22

External Extended Range Fuel System Failure To Transfer Symmetrically In Manual Mode ERFS
External Extended Range Fuel System Kit

External Extended Range Fuel System Kit Configurations

.....................................................................

5.34

External Extended Range Fuel System Tanks ..................................................................................................

4.20.2

..............................................................................

9.36

External Extended Range Fuel Transfer Check................................................................................................

4.20.7

External Extended Range Fuel Transfer In Auto Mode...................................................................................

4.20.7.1

External Extended Range Fuel Transfer In Manual Mode...............................................................................

4.20.7.2

External Extended Range Fuel Transfer Modes ...............................................................................................

4.20.1

External Extended Range Fuel System Tank Jettison

External Load Drag............................................................................................................................................

External Load Drag Chart..................................................................................................................................

7-30

7A-33
7.20
7A.21

External Source Engine Start.............................................................................................................................

2.28.4

External Stores Fixed Provisions.......................................................................................................................

4.21.1

External Stores Jettison Control Panel ..............................................................................................................

4.21.4

External Stores Removable Provisions..............................................................................................................

4.21.2

...................................................................................................

4.23

External Stores Support System (ESSS)

F
Fire......................................................................................................................................................................

9.23

Fire Detection System........................................................................................................................................

2.14.1

Fire Detector Test Panel ....................................................................................................................................

2.14.2

Fire Extinguisher Arming Levers (T-handles) ..................................................................................................

2.14.4

Fire Extinguisher Control Panel ........................................................................................................................

2.14.5

Fire Extinguishing Systems ...............................................................................................................................

2.14.3

Change 5

INDEX-13

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Fire Protection Systems .....................................................................................................................................

2.14

First Aid Kits......................................................................................................................................................

2.16

Flare Dispenser M130

..............................................................................................................................

4.3.2

Flight Characteristics..........................................................................................................................................

9.8

Flight Control/Main-Rotor System Malfunctions .............................................................................................

9.29

Flight Control Servo Low-Pressure Caution Lights .........................................................................................

2.36.2

Flight Control Servo Systems............................................................................................................................

2.36

Flight Control Servo Switch ..............................................................................................................................

2.36.1

Flight Control Systems.......................................................................................................................................

2.35

Flight Data Recorder (On Helicopters Equipped with Flight Data Recorder Kit) ..........................................

2.57

Flight In Icing Conditions..................................................................................................................................

5.28

Flight Instrument Lights.....................................................................................................................................

2.70.3

Flight In Instrument Meteorological Conditions (IMC) ...................................................................................

5.27

FLT PATH STAB Caution Light On................................................................................................................

9.29.4

Flight Path Stabilization (FPS) ..........................................................................................................................

2.37.3

Flight With Cabin Door(s)/Window(s) Open....................................................................................................

5.20

Flight With External Loads ...............................................................................................................................

8.39.1

.........................................................................................

8.39.2

Flying Qualities with External ERFS Installed

FM Control AN/ARC-114A ..............................................................................................................................

3-3

FM Control AN/ARC-201 .................................................................................................................................

3-7

Formation Lights ................................................................................................................................................

2.71.5

Forms and Records.............................................................................................................................................

1.9

Free Air Temperatures .......................................................................................................................................

7.9
7A.9

Free-Air Temperature (FAT) Indicator .............................................................................................................
Fuel and Lubricants, Specifications, and Capacities.........................................................................................

2.79
T

2-4

Fuel Boost Pump................................................................................................................................................

2.34.2

Fuel Filter ...........................................................................................................................................................

2.18.2
2.32.3

......................

9.37

Fuel Limitations .................................................................................................................................................

5.12

Fuel Low Caution Light.....................................................................................................................................

2.34.1

Fuel Fumes in Cockpit/Cabin With External Extended Range Fuel System Pressurized

INDEX-14

Change 5

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Fuel Moments.....................................................................................................................................................

6-2
6.10

Change 5

INDEX-14.1/(INDEX-14.2 Blank)

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Fuel Pressure Warning System..........................................................................................................................

2.18.3

Fuel Quantity Indicating System .......................................................................................................................

2.34

Fuel Sampling System .......................................................................................................................................

2.84.5

Fuel Supply System ...........................................................................................................................................

2.32

Fuel System ........................................................................................................................................................

9.25

Fuel System Servicing .......................................................................................................................................

2.84

Fuel Tanks ..........................................................................................................................................................

2.32.1

Fuel Transfer Sequence......................................................................................................................................

4.20.6

Fuel Types ..........................................................................................................................................................

2.84.1

Fuselage-Left Side (AREA 5) ...........................................................................................................................

8.17

Fuselage-Right Side (AREA 7) .........................................................................................................................

8.19

G
Gas Generator Speed (Ng) Indicator .................................................................................................................
General Arrangement .........................................................................................................................................

2.31.4
F

2-1

Generator Control Switches ...............................................................................................................................

2.66.1

Generator Control Units (GCU) ........................................................................................................................

2.65.1

Gravity Refueling ...............................................................................................................................................

2.84.3

Ground Operations .............................................................................................................................................

8.43.5

Ground Resonance .............................................................................................................................................

8.38

Ground Taxi .......................................................................................................................................................

8.26

Gust Lock Limitations .......................................................................................................................................

5.37

Gyro Magnetic Compass Set AN/ASN-43........................................................................................................

3.24

H
Hand-Operated Fire Extinguishers.....................................................................................................................

2.14.7

Handpump Reservoir Servicing .........................................................................................................................

2.88

Heads Up Display AN/AVS-7...........................................................................................................................

4-7
4.12

Heat and Ventilation Controls ...........................................................................................................................

2.59.2

Heating System...................................................................................................................................................

2.59

............................................................................................................

9-2

....................................................................................................................

9-3

Helicopter Compartment and Station Diagram .................................................................................................

6-1

Height Velocity Diagram

UH−60A

Height Velocity Diagram

UH−60L

Change 1

INDEX-15

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

6.3
Helicopter Systems.............................................................................................................................................

9.1

High Drag Symbol .............................................................................................................................................

1.12

High Speed Yaw Maneuver Limitation.............................................................................................................

5.23.3.1

.......................................................................................................................................

2.22

.......................................................................................................................................

2.21

Horizontal Reference Datum .............................................................................................................................

6.6.1

History Counter

701C

History Recorder

700

Horizontal Situation Indicator............................................................................................................................

3-31
3.25.2

Hover Chart ........................................................................................................................................................

7.14
7A.15

Hover Check.......................................................................................................................................................

8.27

Hover - Clean .....................................................................................................................................................

7A-6

Hover-Clean Configuration ................................................................................................................................

7-4

Hover-High Drag................................................................................................................................................

7-5

7A-7

Hover Infrared Suppressor Subsystem (HIRSS) ...............................................................................................
HSI/VSI MODE SEL Panel

......................................................................................................................

2.30
F

Hydraulic Leak Detection/Isolation System......................................................................................................
Hydraulic Logic Module Operation Principle...................................................................................................

3-21
2.41

2-16

Hydraulic Pump Modules ..................................................................................................................................

2.40

Hydraulic System ...............................................................................................................................................

2.39
9.27

Hydraulic Systems Servicing .............................................................................................................................

2.89

I
Ice and Rain Operation ......................................................................................................................................

8.43.3

Ice Rate Meter Fail or Inaccurate......................................................................................................................

9.34.3

Idle Fuel Flow ....................................................................................................................................................

7-2

7A-2
7.24
7A.25

IFM Amplifier Control.......................................................................................................................................

INDEX-16

Change 1

3-6

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Ignition System...................................................................................................................................................

2.20

Immediate Action Emergency Steps .................................................................................................................

9.2

In-flight ...............................................................................................................................................................

8.43

In-flight Icing......................................................................................................................................................

8.43.4

Increasing % RPM R .........................................................................................................................................

9.15

Index ...................................................................................................................................................................

1.6

Infrared Countermeasure Set AN/ALQ-144(V) ................................................................................................

4.8

Infrared Suppressor System ...............................................................................................................................

7.13
7A.13

Input Module ......................................................................................................................................................

2.45.1

.....................................................................................................................................................

3-24

Installation and Removal of Ammunition Can on Machinegun M60D...........................................................

4-16

Installation of Ejector Control Bag on Machinegun M60D .............................................................................

4-15

Installation of Machinegun M60D on Pintle.....................................................................................................

4-14

INS Page

Instrument Flight ................................................................................................................................................

8.36

Instrument Markings .........................................................................................................................................

5-1

.................................................................................................................................

5-2

...............................................................................................................................

5-3

Instrument Markings

700

Instrument Markings

701C

Instrument Marking Color Codes ......................................................................................................................
Instrument Panel.................................................................................................................................................

5.5
F

2-9
2.10

Instrument Panel

.........................................................................................................................................

2.10.1

Instrument Panel

.........................................................................................................................................

2.10.2

..............................................................

3.18

Integrated Inertial Navigation System (IINS) AN/ASN-132(V)

Intercommunication Control Panel C-6533/ARC .............................................................................................

3-2

Intercommunication System C-6533/ARC ........................................................................................................

3.4

Interior Cabin (AREA 4) ...................................................................................................................................

8.16

Interior Lighting .................................................................................................................................................

2.70

Intermediate and Tail Gear Box Chip/Temperature Systems...........................................................................

2.47.3

Intermediate Gear Box .......................................................................................................................................

2.47.1

J
Jettison Limits ....................................................................................................................................................

5.35

Change 1

INDEX-17

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

L
Landing ...............................................................................................................................................................

8.32

Landing and Ditching.........................................................................................................................................

9.28

Landing Gear Limitations ..................................................................................................................................

5.24

Landing Gear System.........................................................................................................................................

2.9

Landing Light .....................................................................................................................................................

2.71.2

Landing Speed Limitations ................................................................................................................................

5.25

Launcher Racks Jettison .....................................................................................................................

9.38

VOL

LF/ADF Control Panel C-7932/ARN-89...........................................................................................................

3-13

LF/ADF Control Panel AN/ARN-149...............................................................................................................

3-14

Lightning Strike..................................................................................................................................................

9.21

Limitations for Maneuvering With Rescue Hoist Loads..................................................................................

5.23.3.3

Limitations for Maneuvering With Sling Loads ...............................................................................................

5.23.3.2

Limits..................................................................................................................................................................

7.4
7A.4

Litter Support .....................................................................................................................................................

4.24.1

Load Demand System .......................................................................................................................................

2.29.4

Loading Data ......................................................................................................................................................

6.8

Location and Identification of Safety on Machinegun M60D..........................................................................

4-13

Logic Modules....................................................................................................................................................

2.42.4

Loss of Tail Rotor Thrust ..................................................................................................................................

9.22.1

Loss of Tail Rotor Thrust at Low Airspeed/Hover ..........................................................................................

9.22.2

Loss of No. 1 or No. 2 Generator During Blade Deice Operation..................................................................

9.34.4

Lower Console ...................................................................................................................................................

2-8

4-12

M
Machinegun 7.62 Millimeter M60D..................................................................................................................

4.15
Machinegun M60D Installation .........................................................................................................................

Main Landing Gear ...........................................................................................................................................
Main Rotor Blade and BIMt System ...............................................................................................................

4-11
2.9.1

2-17

Main Rotor Blades .............................................................................................................................................

2.49.1

Main Rotor Gust Lock .......................................................................................................................................

2.49.2

INDEX-18

Change 1

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Main Rotor System ............................................................................................................................................

2.49

Main Rotor Tiedown..........................................................................................................................................

2.95.2

Main Transmission BIT Chip Detector .............................................................................................................

2.46.3

Main Transmission Lubrication System............................................................................................................

2.46

Main Transmission Module Limitations ...........................................................................................................

5.7

Main Transmission Oil System Servicing.........................................................................................................

2.91

MAIN XMSN OIL PRESS Caution Light On/XMSN OIL PRESS LOW/XMSN OIL TEMP
HIGH or XMSN OIL TEMP Caution Light On...............................................................................................

9.22.7

Maintenance Light..............................................................................................................................................

2.70.8

Maneuvering Flight ............................................................................................................................................

8.39

Maneuvering Limitations ...................................................................................................................................

5.23.3

Manual Operation of the Stabilator ...................................................................................................................

5.23.1

Master Mode Display.........................................................................................................................................

4-6.2

Master Warning Panel........................................................................................................................................

2-23

Master Warning System.....................................................................................................................................

2.81

Maximum Cargo Size Diagram for Loading Through Cabin Doors ...............................................................

6.17

Maximum Package Size for Cargo Door ..........................................................................................................

6-10

Maximum Torque Available .............................................................................................................................

7A-4

Maximum Torque Available Chart....................................................................................................................

7.11

Maximum Torque Available - 30 Minute Limit...............................................................................................

7-3

Medevac and Seat System .................................................................................................................................

4-29

Medevac Kit Personnel Moments......................................................................................................................

6.12

Medevac Seats Installation.................................................................................................................................

4.25.5

Medical Evacuation (Medevac) Kit...................................................................................................................

4.25

Minimum Crew Requirements...........................................................................................................................

5.4

.....................................................................................................................

3.30

Mission Equipment Interface

Mission Kits .......................................................................................................................................................

4-1

Mission Planning................................................................................................................................................

8.1

Mission Readiness Circuit Breaker Panel .........................................................................................................

4.18

Mixing Unit ........................................................................................................................................................

2.35.3

Moment...............................................................................................................................................................

6.6.3

Mooring ..............................................................................................................................................................

Change 5

2-26

INDEX-19

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

2.95
Mooring Instructions ..........................................................................................................................................

2.95.1

MR DE-ICE FAULT or MR DE-ICE FAIL, or TR DE-ICE FAIL Caution Light On..................................

9.34.1

N
No. 1 Transfer Module ......................................................................................................................................

2.42.1

No. 2 Transfer Module ......................................................................................................................................

2.42.2

Nose Section (AREA 1).....................................................................................................................................

8.13

Number 1 Hydraulic System .............................................................................................................................

2.40.1

Number 2 Hydraulic System .............................................................................................................................

2.40.2

NVG Lighting System .......................................................................................................................................

2.70.1

O
..........................................................................................................................................

2.13.3

Operating Procedures and Maneuvers ...............................................................................................................

8.7

Observer’s Seat

Optimum Altitude For Maximum Range ..........................................................................................................

7-29

Optimum Altitude For Maximum Range ..........................................................................................................

7A-31

Optimum Altitude For Maximum Range - High Drag.....................................................................................

7A-32

Optimum Range Charts......................................................................................................................................

7.19
7A.20

Overspeed and Drain Valve

701C

...................................................................................................................

2.18.4.2

P
Parking................................................................................................................................................................

2.93

Parking and Shutdown .......................................................................................................................................

8.34

Passenger Briefing..............................................................................................................................................

8.6

Pedal Bind/Restriction or Drive With No Accompanying Caution Light .......................................................

9.22.5

Performance Data Basis-Clean ..........................................................................................................................

7.7
7A.7

Performance Data Basis-High Drag ..................................................................................................................

7.8
7A.8

Performance Discrepancies ................................................................................................................................

7.6
7A.6

Personnel Moments ............................................................................................................................................

6-3
6.11

INDEX-20

Change 1

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Personnel Moments (Medevac Configuration)..................................................................................................

6-4

.........................................................................................

3.23

Pilot-Assist Controls...........................................................................................................................................

2.36.5

Pilot-Assist Servos .............................................................................................................................................

2.36.3

Pilot’s Display Unit (PDU)................................................................................................................................

2.10.5

Pilot’s Seats ........................................................................................................................................................

2.12.1

Pitch Boost Servo Hardover ..............................................................................................................................

9.27.10

Pitch, Roll or Yaw/Trim Hardover....................................................................................................................

9.29.5

Pitot Heater.........................................................................................................................................................

2.53

Pitot-Static System .............................................................................................................................................

2.72

Placarded Aircraft Symbol.................................................................................................................................

1.13

Pneumatic Source Inlet Limits...........................................................................................................................

5.9

Pneumatic Subsystem.........................................................................................................................................

2.44

Position Lights....................................................................................................................................................

2.71.4

Pilot and Copilot VSI/HSI MODE SEL Panel

..............................................................................................................................................

3-23

Positioning Cartridge Link Belt on Machinegun M60D ..................................................................................

4-17

Position Page

Powertrain...........................................................................................................................................................

2.45

Preflight Check ...................................................................................................................................................

8.10

Pressure Refueling..............................................................................................................................................

2.84.4

Pressurizing and Overspeed Unit

700

..............................................................................................................

Principal Dimensions .........................................................................................................................................

2.18.4.1
F

2-2

Prohibited Maneuvers.........................................................................................................................................

5.22

Protective Covers and Plugs ..............................................................................................................................

2.94

PWR MAIN RTR and/or TAIL RTR MONITOR Light On ...........................................................................

9.34.2

Q
Quick Fix Power

........................................................................................................................................

2.64.4

R
Radar Altimeter Set AN/APN-209(V)...............................................................................................................

3-34
3.29

Radar Signal Detecting Set AN/APR-39A(V)1 ................................................................................................

4.7

...........................................................................................

4.6

Radar Signal Detector Set AN/APR-39(V)-1 ...................................................................................................

4.4

Change 1

INDEX-21

Radar Signal Detecting Set AN/APR-39(V)2

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Radio Receiving Set AN/ARN-123(V) .............................................................................................................

3-15

Radio Receiving Set AN/ARN-123(V) (VOR/ILS/MB) ..................................................................................

3.15

Radio Receiving Set AN/ARN-147(V) (VOR/ILS/MB) ..................................................................................

3.16

Radio Retransmission Control ...........................................................................................................................

3.11

........................................................................................................

3.5

......................................................................................................

3.6

Radio Set AN/ARC-186(V) ...............................................................................................................................

3.7

Radio Set AN/ARC-114A(VHF-FM)

Radio Set AN/ARC-115A (VHF-AM)

Radio Set AN/ARC-201.....................................................................................................................................
Radio Set, HF AN/ARC-220 .............................................................................................................................

3.12A

Rappeling Rope Connectors...............................................................................................................................

4.24

Receiver-Transmitter Radio, RT-1167/ARC-164(V) ........................................................................................

3.10

Receiver-Transmitter Radio, RT-1167C/ARC-164(V)......................................................................................

3.9

Recommended Emergency External Fuel Tank Jettison Envelope..................................................................

Refueling/Defueling ...........................................................................................................................................

5-1
2.34.3

Remote Fill Panel...............................................................................................................................................

3-11

.......................................................................................................................................

4-25

Rescue Hoist Kit

Rescue Hoist Lubrication System Servicing .....................................................................................................
Rescue Hoist Moments ......................................................................................................................................

2.90
F

6-8

Rescue Hoist System Kit ..................................................................................................................................

4.19

Rescue Hoist Weight Limitations......................................................................................................................

5.18

Reservoir Fill System.........................................................................................................................................

2.43

Restraint Criteria ................................................................................................................................................

6.20.3

Restricted Maneuvers .........................................................................................................................................

5.23

Retransmission Control Panel............................................................................................................................

3-12
3.12

Rotor Blade Deice Kit .......................................................................................................................................

2-18
2.54
4.19

Rotor Limitations ...............................................................................................................................................

5.6

Rotor Speed Limitations ....................................................................................................................................

5.6.2

Rotor Start and Stop Limits...............................................................................................................................

5.6.1

Rotor Systems.....................................................................................................................................................

2.48

INDEX-22

Change 5

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Rotors, Transmissions and Drive Systems ........................................................................................................

Change 5

9.22

INDEX-22.1/(INDEX-22.2 Blank)

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

% RPM Increasing/Decreasing (Oscillation).....................................................................................................

9.16

S
Sample Cruise Chart ..........................................................................................................................................

7A-8

Sample Cruise Chart - Clean .............................................................................................................................

7-6

SAS Failure With No Failure/Advisory Indication...........................................................................................

9.29.1

SAS 2 Failure Advisory Light On.....................................................................................................................

9.29.2

SAS OFF Caution Light On ..............................................................................................................................

9.29.3

Scope...................................................................................................................................................................

6.4

Searchlight ..........................................................................................................................................................

2.71.1

Self Deployment Mission Profile.......................................................................................................................

Self-Test Patterns AN/APR-39(V)2...................................................................................................................

7-38

7A-40

4-6

Series and Effectivity Codes..............................................................................................................................

1.11

Service Platforms and Fairings..........................................................................................................................

2.83

Servicing .............................................................................................................................................................

2.82

Servicing Diagram..............................................................................................................................................
Signal Converter Unit,CV-3739/ASN-132.

...............................................................................................

2-25
3.20

.............................................................................................................

2-12

Single/Dual-Engine Fuel Flow...........................................................................................................................

7-34

7A-37

Signal Validation - Fault Codes

701C

Single-Engine .....................................................................................................................................................

7.18

Single-Engine .....................................................................................................................................................

7A.19

Single-Engine Fuel Flow ...................................................................................................................................

7-25
7A.26

Single-Engine Failure.........................................................................................................................................

9.10

Single-Engine Failure-General...........................................................................................................................

9.9

Sling/Hoist Load Maneuvering Limitations ......................................................................................................

5-10

Slope Landing Limitations.................................................................................................................................

5.26

Smoke and Fume Elimination ...........................................................................................................................

9.24

Special Mission Flight Profiles..........................................................................................................................

7.29
7A.30

Specific Conditions.............................................................................................................................................

7.5.3

Change 1

INDEX-23

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

7A.5.2
Spheroid Data Codes

.................................................................................................................................

3-2

Stabilator Control Panel .....................................................................................................................................

2.38.1

Stabilator Malfunction-Auto Mode Failure .......................................................................................................

9.30

Stabilator Position Indicator...............................................................................................................................

2.38.2

Stabilator System ..............................................................................................................................................

2.38

Stability Augmentation System (SAS) ..............................................................................................................

2.37.1

Standby Magnetic Compass...............................................................................................................................

2.78

Starting Engines .................................................................................................................................................

8.23

.....................................................................................................................

4-28

Stowage Compartment Moments.......................................................................................................................

6-12

Stores Jettison Control Panel

Stowage Provisions ............................................................................................................................................
Symbol Generator Test Mode............................................................................................................................

5.15
F

Symbols Definition.............................................................................................................................................
System Annuciators

...................................................................................................................................

SYSTEMS SELECT Panel

........................................................................................................................

4-9
8.8

3-28

3-20
3.22

T
TACAN Control Page

................................................................................................................................

3-25

..................................................................

3.21

Tail and Intermediate Gear Box Servicing .......................................................................................................

2.92

Tail Drive System ..............................................................................................................................................

2.47

Tail Gear Box.....................................................................................................................................................

2.47.2

Tail Landing Gear ..............................................................................................................................................

2.9.3

Tail Pylon (AREA 6) .........................................................................................................................................

8.18

Tail Rotor Control..............................................................................................................................................

2.35.5

Tail Rotor Pedals................................................................................................................................................

2.35.6

TAIL ROTOR QUADRANT Caution Light On With Loss of Tail Rotor Control ......................................

9.22.4

TAIL ROTOR QUADRANT Caution Light On With No Loss of Tail Rotor Control .................................

9.22.3

Tail Rotor Quadrant/Warning ............................................................................................................................

2.51

Tail Rotor System ..............................................................................................................................................

2.50

Takeoff................................................................................................................................................................

8.29

TACAN Navigational Set Receiver-Transmitter, RT-1159/A

INDEX-24

Change 1

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Taxiing................................................................................................................................................................

8.41.5

Taxiing and Ground Operation..........................................................................................................................

8.42.1

Temperature Conversion Chart..........................................................................................................................

7-1

7A-1

TGT Temperature Indicator ...............................................................................................................................

2.31.3

The Main Module...............................................................................................................................................

2.45.3

Thermocouple Harness.......................................................................................................................................

2.23

Thunderstorm Operation ....................................................................................................................................

8.43.1

Tiedown Fittings and Restraint Rings ...............................................................................................................

6.18

Torque and Overspeed and % RPM Sensors....................................................................................................

2.24

Torque Available................................................................................................................................................

7A.11

Torque Available - 10 Minutes .........................................................................................................................

7A.11.1

Torque Available - 30 Minutes .........................................................................................................................

7A.11.2

Torque Conversion Chart...................................................................................................................................
Torque Factor Method .......................................................................................................................................

7A-3
7.10
7A.10

Torque Factor Procedure....................................................................................................................................

7.10.2
7A.10.2

Torque Factor Terms..........................................................................................................................................

7.10.1
7A.10.1

Torque Indicator .................................................................................................................................................

2.31.6

Transfer Modules ...............................................................................................................................................

2.42

Transmission Chip Detector System .................................................................................................................

2.46.4

Transmission Oil Pressure Indicator..................................................................................................................

2.46.2

Transmission Oil Temperature Indicator...........................................................................................................

2.46.1

Transponder AN/APX-100(V)1(IFF).................................................................................................................

3.26

Transponder Computer KIT-1A/TSEC..............................................................................................................

3.27

Trim Actuator Jammed ......................................................................................................................................

9.29.6

Trim System .......................................................................................................................................................

2.37.2

Troop/Cargo (Cabin) Doors ...............................................................................................................................

2.11.2

Troop Commander’s Antenna............................................................................................................................

4.1

........................................................................................................................................

2.13

Change 1

INDEX-25

Troop Provisions

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Troop Seat Belt Operation
Troop Seats

.........................................................................................................................

.................................................................................................................................................

2.13.1
F

2-10

% TRQ Split Between Engines 1 and 2 ...........................................................................................................

9.17

Turbulence ..........................................................................................................................................................

8.43.2

Turbulence and Thunderstorm Operation..........................................................................................................

5.33

Turn Rate Indicating System .............................................................................................................................

2.74

Turning Radius and Clearance...........................................................................................................................

Turning Radius and Ground Clearance .............................................................................................................
Typical High Drag Configurations ....................................................................................................................

Typical Self-Test Mode Display........................................................................................................................

2-3
2.6

7-31

7A-34

4-4

U
UH-60A

UH−60A

................................................................................................................................................

UH-60A/L Master Mode Symbology Display (HUD) .....................................................................................

UH-60L

UH−60L

2.2
T

4-1

4-8

................................................................................................................................................

UHF Control, AN/ARC-164(V) ........................................................................................................................

2.3
F

Uncommanded Nose Down/Up Pitch Attitude Change....................................................................................

3-8
9.31

...............................................................................................................................................

3-26

Upper Console....................................................................................................................................................

2-7

Update Page

Upper and Lower Console Lights .....................................................................................................................

2.70.5

Upper and Lower Consoles ...............................................................................................................................

2.8

Use of Charts......................................................................................................................................................

7.5
7.17
7A.5
7A.18

Use of Fuels .......................................................................................................................................................

2.84.2

........................................................................................

5.36

Use of Words Shall, Should, and May .............................................................................................................

1.14

Utility Lights ......................................................................................................................................................

2.70.6

Utility Module ....................................................................................................................................................

2.42.3

Use of M60D Gun(s) with ERFS Kit Installed

INDEX-26

Change 1

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

V
Ventilation System .............................................................................................................................................

2.60

Ventilation System

.....................................................................................................................................

2.60.2

Ventilation System

.....................................................................................................................................

2.60.1

Vertical Instrument Display System (VIDS).....................................................................................................

2.10.3

Vertical Situation Indicator................................................................................................................................

3-30
3.25.1

Vertical Speed Indicator.....................................................................................................................................

2.77

VHF Control AN/ARC-115A ............................................................................................................................

3-4

VHF Control Panel AN/ARC-186(V) ...............................................................................................................

3-5

Voice Security Equipment .................................................................................................................................

3-10

Voice Security System .......................................................................................................................................

3.11

Voice Security System Control C-8157/ARC...................................................................................................

3-9

Volcano ICP .......................................................................................................................................................

4-22

Volcano Mine Dispenser Controls.....................................................................................................................

4-19

Volcano Mine Moments.....................................................................................................................................

6-6

Volcano Multiple Mine Delivery System .........................................................................................................

4.16

Volcano System DCU Displays.........................................................................................................................

4-20

VOR/ILS/MB Control Panel AN/ARN-147 (V)...............................................................................................

3-16

VSI/HSI and CIS Mode Selector Panels...........................................................................................................

3.25.4

W
Warnings, Cautions, and Notes .........................................................................................................................

1.2

Weight Definitions .............................................................................................................................................

6.5

Weight Limitations.............................................................................................................................................

5.14

Weight-On-Wheels Functions............................................................................................................................

2-1

Wheel Brake System..........................................................................................................................................

2.9.2

Windshield Anti-Ice/Defogging System............................................................................................................

2.52.2

Windshield Anti-Ice Limitations .......................................................................................................................

5.32

Windshield Wiper Control .................................................................................................................................

2.52.1

Windshield Wipers .............................................................................................................................................

2.52

Winterized Heater...............................................................................................................................................

2.59.1

Change 1

INDEX-27

TM 1-1520-237-10

INDEX (Cont)
Subject

Paragraph
Figure, Table
Number

Wire Strike Protection System...........................................................................................................................

INDEX-28

Change 1

2.56

TM 1-1520-237-10

By Order of the Secretary of the Army:

DENNIS J. REIMER
General, United States Army
Chief of Staff

Official:

JOEL B. HUDSON
Administrative Assistant to the
Secretary of the Army
03477

DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E,
block no. 1284, requirements for TM 1-1520-237-10.

RECOMMENDED CHANGES TO EQUIPMENT TECHNICAL PUBLICATIONS

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PUBLICATION NUMBER

TM 1-1520-237-10

WITH THIS PUBLICATION?

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DATE SENT

PUBLICATION DATE

30 December 1998

10 January 1999
PUBLICATION TITLE

Operator's Manual for UH-60

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DEPARTMENT OF THE ARMY

OFFICIAL BUSINESS

COMMANDER
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ATTN: AMSAM-MMC-MA-NP
REDSTONE ARSENAL, AL 35898-5230

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FILL IN YOUR
UNITS ADDRESS

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THEN . .JOT DOWN THE
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FORM, CAREFULLY TEAR
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1 JUL 79

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REVERSE OF DA FORM 2028-2 Reverse of DRSTS-M Overprint 2,
1 Nov 80

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DEPARTMENT OF THE ARMY

OFFICIAL BUSINESS

COMMANDER
U.S. ARMY AVIATION AND MISSILE COMMAND
ATTN: AMSAM-MMC-MA-NP
REDSTONE ARSENAL, AL 35898-5230

TEAR ALONG PERFORATED LINE

FILL IN YOUR
UNITS ADDRESS

RECOMMENDED CHANGES TO EQUIPMENT TECHNICAL PUBLICATIONS

SOMETHING WRONG
THEN . .JOT DOWN THE
DOPE ABOUT IT ON THIS
FORM, CAREFULLY TEAR
IT OUT, FOLD IT AND
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FROM: (PRINT YOUR UNIT’S COMPLETE ADDRESS)

DATE SENT

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PUBLICATION TITLE

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PIN--POINT WHERE IT IS IN THIS SPACE, TELL WHAT IS WRONG
AND WHAT SHOULD BE DONE ABOUT IT:
PAGE PARA- FIGURE TABLE
NO
NO
GRAPH
NO

PRINTED NAME, GRADE OR TITLE, AND TELEPHONE NUMBER

FORM

1 JUL 79

2028--2

PREVIOUS EDITIONS
ARE OBSOLETE.
DRSTS-M verprint2, 1 Nov 80

SIGN HERE

P.S.- - IF YOUR OUTFIT WANTS TO KNOW AB OUT Y OUR
RECOMMENDATION, MAKE A CARBON COPY OF THIS
AND GIVE TO YOUR HEADQUARTERS.

REVERSE OF DA FORM 2028-2 Reverse of DRSTS-M Overprint 2,
1 Nov 80

FOLD BACK

DEPARTMENT OF THE ARMY

OFFICIAL BUSINESS

COMMANDER
U.S. ARMY AVIATION AND MISSILE COMMAND
ATTN: AMSAM-MMC-MA-NP
REDSTONE ARSENAL, AL 35898-5230

TEAR ALONG PERFORATED LINE

FILL IN YOUR
UNITS ADDRESS

The Metric System and Equivalents
Linear Measure

Liquid Measure
1 centiliter = 10 milliters = .34 fl. ounce
1 deciliter = 10 centiliters = 3.38 fl. ounces
1 liter = 10 deciliters = 33.81 fl. ounces
1 dekaliter = 10 liters = 2.64 gallons
1 hectoliter = 10 dekaliters = 26.42 gallons
1 kiloliter = 10 hectoliters = 264.18 gallons

1 centimeter = 10 millimeters = .39 inch
1 decimeter = 10 centimeters = 3.94 inches
1 meter = 10 decimeters = 39.37 inches
1 dekameter = 10 meters = 32.8 feet
1 hectometer = 10 dekameters = 328.08 feet
1 kilometer = 10 hectometers = 3,280.8 feet

Square Measure
Weights
1 sq. centimeter = 100 sq. millimeters = .155 sq. inch
1 sq. decimeter = 100 sq. centimeters = 15.5 sq. inches
1 sq. meter (centare) = 100 sq. decimeters = 10.76 sq. feet
1 sq. dekameter (are) = 100 sq. meters = 1,076.4 sq. feet
1 sq. hectometer (hectare) = 100 sq. dekameters = 2.47 acres
1 sq. kilometer = 100 sq. hectometers = .386 sq. mile

1 centigram = 10 milligrams = .15 grain
1 decigram = 10 centigrams = 1.54 grains
1 gram = 10 decigram = .035 ounce
1 decagram = 10 grams = .35 ounce
1 hectogram = 10 decagrams = 3.52 ounces
1 kilogram = 10 hectograms = 2.2 pounds
1 quintal = 100 kilograms = 220.46 pounds
1 metric ton = 10 quintals = 1.1 short tons

Cubic Measure
1 cu. centimeter = 1000 cu. millimeters = .06 cu. inch
1 cu. decimeter = 1000 cu. centimeters = 61.02 cu. inches
1 cu. meter = 1000 cu. decimeters = 35.31 cu. feet

Approximate Conversion Factors
To change

inches
feet
yards
miles
square inches
square feet
square yards
square miles
acres
cubic feet
cubic yards
fluid ounces
pints
quarts
gallons
ounces
pounds
short tons
pound-feet
pound-inches

centimeters
meters
meters
kilometers
square centimeters
square meters
square meters
square kilometers
square hectometers
cubic meters
cubic meters
milliliters
liters
liters
liters
grams
kilograms
metric tons
Newton-meters
Newton-meters

Multiply by

To change

2.540
.305
.914
1.609
6.451
.093
.836
2.590
.405
.028
.765
29,573
.473
.946
3.785
28.349
.454
.907
1.356
.11296

ounce-inches
centimeters
meters
meters
kilometers
square centimeters
square meters
square meters
square kilometers
square hectometers
cubic meters
cubic meters
milliliters
liters
liters
liters
grams
kilograms
metric tons

To
Newton-meters
inches
feet
yards
miles
square inches
square feet
square yards
square miles
acres
cubic feet
cubic yards
fluid ounces
pints
quarts
gallons
ounces
pounds
short tons

Temperature (Exact)
°F

Fahrenheit
temperature

5/9 (after
subtracting 32)

Celsius
temperature

°C

Multiply by
.007062
.394
3.280
1.094
.621
.155
10.764
1.196
.386
2.471
35.315
1.308
.034
2.113
1.057
.264
.035
2.205
1.102

PIN: 073161-010

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Operator's Manual For UH 60A Helicopter TM 1 1520 237 10

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