D2007 3 13 S 206 & T206 SERIES (1969 THRU 1976) Cessna_206_T206_1969 1976_MM_D2007 Cessna 1969 1976 MM

User Manual: Cessna_206_T206_1969-1976_MM_D2007-3-13

Open the PDF directly: View PDF .
Page Count: 542

Download
Open PDF In Browser	View PDF

Cessna

SERVICE MANUAL

1969
thru
1976

&

MODEL 206
T206 SERIES
Member of

THIS

REPRINT

GAMA

OF BASIC SERVICE MANUAL D2007-13, DATED 15 OCTOBER

1972, INCORPORATES CHANGE 1, DATED 15 OCTOBER 1973; CHANGE 2, DATED 1
SEPTEMBER 1974; CHANGE 3, DATED 1 OCTOBER 1975; TEMPORARY CHANGE 1,
DATED 5 SEPTEMBER 1977; AND TEMPORARY CHANGE 2, DATED 22 JANUARY 1978.

)

COPYRIGHT © 1984

CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS. USA

D2007-3-13
(RGI-100-10/01)

15 OCTOBER
15 OCTOB

CHANGED

1972
1972

1 OCTOBER 1975

A Te~t

Gr

-y(~''n

TEMPORARY REVISION NUMBER 7
DATE July 1, 2007
MANUAL TITLE

Model 206 and T206 (1969-1976) Service Manual

MANUAL NUMBER

-

MANUAL NUMBER

-AEROFICHE

PAPER COPY

D2007-3-1 3
D2007-3-1 3AF

TEMPORARY REVISION NUMBER
MANUAL DATE

D2007-3TR7

15 October 1972

REVISION NUMBER 3

DATE

1 October 1975

This Temporary Revision consists of the following pages, which affect and replace existing pages
in the paper copy manual and supersede aerofiche and CD information.
SECTION
5
5
5
5

PAGE
4A
4A 1
4A2
4A3

AEROFICHE
FICHE/FRAME

SECTION

PAGE

AEROFICHE
FICHE/FRAME

1D19
ADD
ADD
ADD

REASON FOR TEMPORARY REVISION
1. Incorporated inspection of flat spring main landing gear (Section 5).
FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
1. For Paper Publications, file this cover sheet behind the publication's title page to identify the
inclusion of the Temporary Revision into the manual. Insert the new pages into the publication
at the appropriate locations and remove and discard the superseded pages.
2. For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche
frame (page) affected by the Temporary Revision. This will be a visual identifier that the
information on the frame (page) is no longer valid and the Temporary Revision should be
referenced. For "added" pages in a Temporary Revision, draw a vertical line between the
applicable frames. Line should be wide enough to show on the edges of the pages. Temporary
Revisions should be collected and maintained in a notebook or binder near the aerofiche library
for quick reference.
3. For CD publications, mark the temporary revision part number on the CD label with permanent
red marker. This will be a visual identifier that the temporary revision must be referenced when
the content of the CD is being used. Temporary revisions should be collected and maintained in
a notebook or binder near the CD library for quick reference.

©9
CESSNA

AIRCRAFT COMPANY

Cessna

A Texlrn Company

TEMPORARY REVISION NUMBER 6
DATE 5 April 2004
MANUAL TITLE

Model 206 & T206 Series 1969 Thru 1976 Service Manual

MANUAL NUMBER - PAPER COPY

D2007-3-13

MANUAL NUMBER - AEROFICHE

D2007-3-13AF

TEMPORARY REVISION NUMBER

D2007-3TR6

MANUAL DATE 15 October 1972

REVISION NUMBER.

3

DATE

1 October 1975

This Temporary Revision consists of the following pages, which affect and replace existing pages
in the paper copy manual and supersede aerofiche information.

SECTION

PAGE

2
2

24
27

FICHE/FRAME

SECTION

PAGE

FICHE/FRAME

1/B12
1/B15

REASON FOR TEMPORARY REVISION
1. To add the cleaning interval of the engine fuel injection nozzles.
FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
1. For Paper Publications, file this cover sheet behind the publication's title page to identify the
inclusion of the Temporary Revision into the manual. Insert the new pages into the publication
at the appropriate locations and remove and discard the superseded pages.
2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche
frame (page) affected by the Temporary Revision. This will be a visual identifier that the
information on the frame (page) is no longer valid and the Temporary Revision should be
referenced. For "added" pages in a Temporary Revision, draw a vertical line between the
applicable frames. Line should be wide enough to show on the edges of the pages. Temporary
Revisions should be collected and maintained in a notebook or binder near the aerofiche library
for quick reference.

© Cessna Aircraft Company

Cessn

A Textron Company

TEMPORARY REVISION NUMBER 5
DATE 6 January 2003
MANUAL TITLE

Model 206 & T206 Series 1969 Thru 1976 Service Manual

MANUAL NUMBER - PAPER COPY

D2007-3-13

MANUAL NUMBER - AEROFICHE

D2007-3-13AF

TEMPORARY REVISION NUMBER

D2007-3TR5

MANUAL DATE 15 October 1972

REVISION NUMBER

3

DATE

1 October 1975

This Temporary Revision consists of the following pages, which affect and replace existing pages
in the paper copy manual and supersede aerofiche information.
SECTION

PAGE

FICHE/FRAME

2
2
2
2
2
2
2
2
2
16
16

24
24A/Delete
25
26
26A/Delete
27
28
29
30
18C
18D

1/B12
N/A
1/B13
1/B14
N/A
1/B15
Added
Added
Added
Added
Added

SECTION

PAGE

FICHE/FRAME

REASON FOR TEMPORARY REVISION
1. To add a Component Time Limits section and a fuel quantity indicating system operational

test.

FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
1. For Paper Publications, file this cover sheet behind the publication's title page to identify the
inclusion of the Temporary Revision into the manual. Insert the new pages into the publication
at the appropriate locations and remove and discard the superseded pages.
2. For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche
frame (page) affected by the Temporary Revision. This will be a visual identifier that the
information on the frame (page) is no longer valid and the Temporary Revision should be
referenced. For "added" pages in a Temporary Revision, draw a vertical line between the
applicable frames. Line should be wide enough to show on the edges of the pages. Temporary
Revisions should be collected and maintained in a notebook or binder near the aerofiche library
for quick reference.

COPYRIGHT © 2003
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA

TEMPORARY REVISION NUMBER 4
DATED 15 May 2000
MANUAL TITLE

MODEL 206 & T206 SERIES 1969 THRU 1976 SERVICE MANUAL

MANUAL NUMBER - PAPER COPY D2007-3-13

AEROFICHE

TEMPORARY REVISION NUMBER PAPER COPY D2007-3TR4
MANUAL DATE

15 OCTOBER 1972 REVISION NUMBER

D2007-3-13AF
AEROFICHE N/A

3 DATE 1 OCTOBER 1975

This Temporary Revision consists of the following pages, which affect existing pages in the
paper copy manual and supersede aerofiche information.
SECTION
2
2

PAGE
24A
26A

AEROFICHE
FICHE/FRAME

SECTION

PAGE

AEROFICHE
FICHE/FRAME

Added
Added

REASON FOR TEMPORARY REVISION
To include the inspection requirements of Cessna Service Bulletin SEB99-18.
FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
For Paper Publications:
File this cover sheet behind the publication's title page to identify the inclusion of the
Temporary Revision into the manual. Insert the new pages into the publication at the
appropriate locations. Draw a line, with a permanent red ink marker, through any
superceded information.
For Aerofiche Publications:
Draw a line through any aerofiche frame (page) affected by the Temporary Revision with a
permanent red ink marker. This will be a visual identifier that the information on the frame
(page) is no longer valid and the Temporary Revision should be referenced. For "added"
pages in a Temporary Revision, draw a vertical line between the applicable frames which is
wide enough to show on the edges of the pages. Temporary Revisions should be collected
and maintained in a notebook or binder near the aerofiche library for quick reference.

COPYRIGHT © 2000
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA

TEMPORARY REVISION NUMBER 3
DATED
MANUAL TITLE

MODEL 206 & T206 SERIES 1969 THRU 1976 SERVICE MANUAL

MANUAL NUMBER - PAPER COPY

D2007-3-13

TEMPORARY REVISION NUMBER - PAPER COPY
MANUAL DATE

3 October 1994

15 OCTOBER 1972

AEROFICHE
D2007-3TR3-13

REVISION NUMBER

3

D2007-3-13AF
AEROFICHE
DATE

N/A

1 OCTOBER 1975

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy
manual and supersede aerofiche information.

SECTION
16
16
16
16
16

PAGE

AEROFICHE
FICHE/FRAME

17
18
18A
18B
18C/D

SECTION

PAGE

AEROFICHE
FICHE/FRAME

2 C21
2 C22
2 C23
added
added

REASON FOR TEMPORARY REVISION
1. To revise procedure to incorporate both Stewart Warner and Rochester fuel gage transmitter calibration.
2. To revise procedures to incorporate both electrically and pressure controlled oil temperature.
3. To add tables to aid in trouble shooting the cylinder head and oil temperature gages.
FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
For Paper Publications:
File this cover sheet behind the publication's title page to identify the inclusion of the Temporary Revision into
the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the
superseded pages.
For Aerofiche Publications:
Draw a line through any aerofiche frame (page) affected by the Temporary Revision with a permanent red ink
marker. This will be a visual identifier that the information on the frame (page) is no longer valid and the
Temporary Revision should be referenced. For "added" pages in a Temporary Revision, draw a vertical line
between the applicable frames which is wide enough to show on the edges of the pages. Temporary Revisions
should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

COPYRIGHT © 1994
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA

INSERT LATEST CHANGED

LIST OF EFFECTIVE PAGES

PAGES.

DESTROY SUPERSEDED

PAGES.

The portion of the text afFcted by Lhe cange is indicated by a vertical
INOTE:
line In the outer marl
o the papg. Change. to illustraUon are indicated by
of
miniature pointinglrhands.
areaU.

Changes to

Dates of issue for original and changed pages are:
Change . .. 2
Original . . . 0 . . 15 October 1972
Change . . 3
15 October 1973
Change . . . . .

iring dialan

are indicated by shaded

1 September 1974
1 October 1975

TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 556, CONSISTING OF THE FOLLOWING:
Page

Change

Page

No.

No.

No.

'Title

......
.
A .
. .
.....
I thru i*5-SB
i ....
.
iLl.
.......
iv Blank .....
1-1 . . . . . . .
1-2 . .......
I-3 .......
1-4
.....
...
1-5 .
....
.2
1-6 Blank .
..'2-1 .......
..
2-2 ........
'2-3 thru 2-4 .- .
2-5 ........
2-6
.. .....
..
2-7 .........
..
2-8
.....
· 2-9
........
2-10 thru 2-11 . ..
'2-12 thru 2-14
2-15 thru 2-17 .
2-18
. .
- '2-19 .
......
2-20 thru 2-25..
2-26 ..
2-27 .
.....
. .
'2-26 Blank .
..
3-1
. . . . . . . .2
3-2 .
..
'3-3 thru 3-6 ...
'3-6A
......
.
3-6B
. ....
.
'3-7 thru 3-8
3-8A .
.
.
3-8 Blank ..
3-9
. . . . . .
3-10 thru 3-11 3-12
.
3-13 thru 3-14
3-14A .
.
3-14B Blank.
.
3-15 thru 3-20
3-213-22 .......
.....
· 3-22.A
.
.
3-22B.
.
3-23
3-24 ......
·
3-25 ·
. . .0
3-26 thri 3-29
3-30 Blank
4-1 thru 4-2
4-3
.
. .
4-4 thru 4-6
. . . .
4- . . . . . . .
4-8 Blank ....
:5-1 . .
5-2 .
-..
*5-3 thru 5-4 .
'5-4A thru 5r4C . .'5-4D Blank .....
5-5.
.
.
.
5-6 .........
5- . . . . . . . . .2

Chge Page

3 * 5-8
......
..
3 '5-8A ...
3
Blank ..
..
0 '5-9 thru 4-13
.
0
5-14 thin 5-17 ....
0
5-18
. . . . . .
2
5-19 thru 5-21 ..
..
3
5-22.
.
........
0 *5-2ZA thru 5-22B.
5-23 thn 5-24 ....
2 *6-1 .........
3
-2 .0
2
6-3 ...
3
6-4
.........
0
6-5 .........
3
6-6
.. .
2
6-7 .........
0
6.......
.
3
7-1
.
.
0
7-2 .0
3
7-3
.
......
I-6 7?-4 thru
0
7-7 thru -8 .....
3
7-9
.
..
1
7-10 thun 7-13 . ...
2 '7-IZA ...
3
7-12B Blink ...
..
3
7-13.
......
*7-14.
.
....
0
8..
.
3
8-2 thru 8-7 .....
3 '8-8 thru 8-9 . ...
2
-10 Blink
3
9-1 thru 9-2......
.....
2
9-2A thrun 9-28
. .
2
9-3 thru 9-4.....
3
9-4A ......
1 9-4B Blaink..3
9-5 .
..
.
. . .
1 '9-6
.
.1
3
9-7 thru 9-12
...
3
9-I2A .
.3
1 '9-12.
.
0
9-13 .
.
3
9-4 thru 9-15
3
9-1 Blank .....
10-1
.
.
2
10-2 .
..
0
.
10-3 .
3
10.4 .
.
.
......
'10-5
10-6
.
0
3
10-7 .
1I 10-8.-1
3
10-9 . ....
I
10-10 Blank
.
1
11-1 thru 11-4
.
3 '12-1 thru 12-4
. . 0
12-5......
3
12-6. . .. .16-1
3
12-7..
.. . .
3
12-8 thru 12-10 .
.
2 '111 .......
. .
3
12-12 thru 12-1
12-14 thru 12-S . .

No.

No.

Change Page
No.

No.

Change Page
No.

No.

A

Change 3

No.

3
12-18 thru 12-17. .
. 2
1
.
.......
20-13
....
3118
.
.
..
. .3
...
20-14 ..
3
12-19 thru 12-20 .
. 0
16-8 .
.......
1 20-1 thru 20-16
3
......
12-21
2-16-9
.
......
.
0 20-17 .
.
.
3
0
12-23.
16-10 ........
........
I
0-18 thru 20-1
.
12-23
.... . .
.
0 '16-11 thnI 16-12 . . . 3 20-20 . . . . . . .
0
12-24 .....
.16-12A
I
Lhru 16-12B 3 20-21 thru 20-25 .
1
12-25 ....
..
3
16-13 thni 16-14
1 20-26 .......
3
12-26 thru 12-27 ...
0
16-14A
......
20-26A ......
0
12-28
......
I15-14B Blank ....
20-268 Blank. . .
2
3
12-29 ........
I
16-15 thru 16-17 20-thru 20-28 . . 1
12-30
...
0
16-18 ...
...
. .
20-29 .
.......
3
I
12-31 thru 12-34
· I
16-8A
. . .
I 20-30 thnu20-32 .
2
1-35 ......
. 3
16-18B Blank. . .*20-33
thin 20-36 .
3
3
12-36 .....
th
1-9
1-20
1 20-37 th 20-42 .
I
12A3
18-21 th
15-22 .
- 0 20-43 ...
2
0
12A-2.
16-23 thr 1-24 .
. 120-44 .......
2
20-44A .....
IIA-3 . .......
2 17- thru 1-2 ....
3
I.I.
.th
1
2-8.A-4
thru 172A-8l
I 204 Blank. . . . 2
12A
..... .
. I
17-A .
. . .
20-45 thru 20-47 .
1
3
12A-8B Blaik.
I
17-6B . .. . . .
'20-48 thru 20-51 .
3
I
12-9
1-17-7 th
thr 1-11
1 11 ......
20-52 thru 20-53
I.
0
12-10 thru 1IA-13
0 1-12 .....
.
20-5 .......
3
. '12A-14 thru 12A-15
. 3
17-13 .
.....
1 20-35 th
0-5 .
1
12A-& .....
0
1-14 thin 17-16
- - 20-57 .......
3
12AI-7 thA 12k1-I.
. I
17-17 .......
.
20-58 thrui 20-59 . - 1
1A-19 .
. .. .
0
17-18 ......
. . . 2 20-60 thru
20-61 .
3
.12A-20
.
. ..
.
3 17-19
......
.
20-62 thru 20-63 .
1
. .. ...I-20
1 th
7· 24
. 3 '3*20-64
.
....
..
3
313A-22 thru 12A-27 . . 0 '17-24A .....
3 20-65 thru 20-68 .
1
IA2....
.
1 '17-24B Blank . .
3 20-69 thru 20-70 .
2
3
12A-29. ....
.
3 17-25 thin 17-26 ·
0 20-70A thnI 20-70E 3
LAAI.k 3 .
0 '17-11 Ithr 11-28
3 ·20-'1 thru 20-73
3.
13-1
3
17-29
.. . .
20-74 thru 20-76 .
13-2 .........
2
17-30 thru 17-32 - 1 20-77 ..
.3
1
. .
20-78
.
.
13-3 ........
0
17-33 thniru 17-34
1
3
13-4 ........
2 17-35
......
.
20-79 . . . . . . .3
0 . 1-8-36 .0......
. .
20-80 .......
.th
'13-9.
..... .
.
.
3 1-37 thru 17-38
·.
2 '20-81 thru 20-82 .
3
20-83 . . . .
.
13-11 . Deleted 17-39 . . . . . .
3
13-10 th
thru 20-86 . . 2
. 20-84
.
17-40 thru 17-42 ·.
13-12 ......
20-86A.
. . 2
2 17-42A. ......
13-13 thrn 13-14 .
- 2
220-868 Blank .
17-42B Blank . .
.
.. .13-15
1
20-87 thru 20-88 .
0
1'7-43 ......
133-6· th r
17 . ·
3
.
20-89 . . . .
3
13-18 thru 13-20 . . · 3 17-44 thru 17-45 .
.....
1
'I
17-46
thru
17-52
.
3
20-90
3
14-1 thru 14-2
. . . .3
2'20-91 .
0
17-53.
. . . ..
.......
i
14-3
1
2
0-92 .......
17-54 Blank ..
i
0
14-4 ........
20-93 ...
. . . .3
0
-1 ........
1-5
14.....
........
3 20-94 .....
14-6.
. .
I
18-2 thru 18-5 ..
3
.
20-94A .
.. . . . .
18-6
. 0
14-7 .......
3
320-94B Blank . .
18-7.
. .....
.
0
14-8 thru 14-10.
15-1 ........
0
18-8 th
18-28. . . . 0 20-95 thru 20-99 .
I
.
. . .
'18-29 ......
..
20-100 thru 20-101
3
I
'15-2A ........
3
18-30 ......
.
20-102 thin 20-103
1
'15-2B Blnk . ...
3
18-31 . .....
'20-104. .
3
1-3 . .....
0
18-32 Blank
..
... 2
0-105 thru 20-106 .
3
3 20-107 .....
15-4 thru 15-5 .
3
19-1 thn 19-2 .
15-6 thru 15-12.
. . 0 20-1 thn 20-3 . . .
3 20-108. . .
. .1
. . .......
1 20-4 thiru 20-5 ..
0 '20-109 thru 20-110
3
16-2.
........
2 *20-6 ........
20-111.
..
I
2
16-3.
.
. . I
0-7 thru 20-10.
. . 1 '20-112 . ....
.
16-4 .....
..
0 '20-11 thru 20-12
3 20-113 thn. 20-115
I
1
.....
0-I.
320-11 .
...
2
0
20-12 Blank . .
.
3 20-116 . ..
.
2
2
20-116B Blank
*20-117 Ir, 20-1
20-I8
3

Upon receipt of the second and subsequent changes to this book, personnel
responsible for maintaining this publication in current status should ascertain
that all previous changes have been received and incorporated.
'The

Change

asterisk indicates pages changed, added or deleted by the current change.

OF CONTENTS

TABLE

Page

SECTION
1

GENERAL DESCRIPTION ........................

1-1

2

GROUND HANDLING, SERVICING, CLEANING, LUBRICATION AND
............
INSPECTION ................

2-1

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3-1

3

FUSELAGE

4

WINGS AND EMPENNAGE

5

LANDING GEAR AND BRAKES

6

AILERON CONTROL SYSTEM ......................

6-1

7

WING FLAP CONTROL SYSTEM .....................

7-1

8

ELEVATOR CONTROL SYSTEMS

9

ELEVATOR TRIM CONTROL SYSTEM

8-1

....................

9-1

..................

10-1

......................

RUDDER CONTROL SYSTEM

11

RUDDER TRIM CONTROL SYSTEM

12

NORMALLY ASPIRATED ENGINE ....................
TURBOCHARGED ENGINE

5-1

.....................

10

12A

4-1

.......................

11-1

...................

12-1
12A-1

.......................

13-1

............................

13

FUEL SYSTEM.

14

PROPELLERS AND PROPELLER GOVERNORS

15

UTILITY SYSTEMS ...........................

15-1

16

INSTRUMENTS AND INSTRUMENT SYSTEMS ...............

16-1

17

ELECTRICAL SYSTEMS

18

STRUCTURAL REPAIR .....................

19

PAINTING .......

20

WIRING DIAGRAMS

14-1

.........

17-1

........................

............
..

....

......................

.

18-1
19-1

............
.

20-1

Change 3

CROSS REFERENCE LISTING OF POPULAR
NAME VS. MODEL NUMBERS AND SERIALS
All aircraft, regardless of manufacturer, are certificated under model number designations. However, popular
names are often used for marketing purposes. To provide a consistent method of referring to the various aircraft. model numbers will be used in this publication unless names are required to differentiate between versions
of the same basic model. The following table provides a cross reference listing popular name vs. model number.
POPULAR NAME
SKYWAGON 206

MODEL
YEAR

MODEL

BEGINNING

ENDING

1969

U206D

U206-1235

U206-1444

P206-0520

P206-0603

U20601445

U20601587

TURBO SKYWAGON 206
SUPER SKYLANE

TU206D
1969

TURBO-SYSTEM
SUPER SKYLANE
SKYWAGON 206

P206D
TP206D

1970

TURBO SKYWAGON 206
SUPER SKYLANE

SERIALS

U206E
TU206E

1970

P206E

P20600604

P20600647

1971

U206E

U20601588

U20601700

1972

U206F

U20601701

U20601874

1973

U206F

U20601875

U20602199

1974

U206F

U20602200

U20602579

1975

U206F

U20602580

U20603020

1976

U206F

U20603021

TURBO SUPER SKYLANE
STATIONAIR
TURBO STATIONAIR
STATIONAIR
TURBO STATIONAIR
STATIONAIR
TURBO STATIONAIR
STATIONAIR
TURBO STATIONAIR
STATIONAIR
STATIONAIR II
TURBO STATIONAIR
TURBO STATIONAIR II
STATIONAIR
STATIONAIR II
TURBO STATIONAmR
TURBO STATIONAIR 11
ii

~~~~ii
3 Change
Change 3

FOREWORD

This manual contains factory recommended procedures and instructions for ground handling, servicing and maintaining Cessna
Stationair, Skywagon and Super Skylane 206-Series aircraft. Also
included are the turbocharged versions of these aircraft.
In addition to this book serving as a reference for the experienced mechanic, it also covers step-by-step procedures for the
less experienced man. This manual should be kept in a handy place
for ready reference. If properly used, it will better enable the
mechanic to maintain Cessna 206 Series aircraft and thereby
establish a reputation for reliable service.
The information in this book is based on data available at the
time for publication, and is supplemented and kept current by service
letters and service news letters published by Cessna Aircraft Company. These are sent to all Cessna Dealers so that they have the
latest authoritative recommendations for servicing Cessna aircraft.
Therefore, it is recommended that Cessna owners utilize the knowledge and experience of the factory-trained Dealer Service Organization.
In addition to the information in this Service Manual, a group
of vendor publications is available from the Cessna Service Parts
Center which describe complete disassembly, overhaul, and parts
breakdown of some of the various vendor equipment items. A listing of the available publications is issued periodically in service
letters.
Information for Nav-O-Matic Autopilots, Electronic Communications and Navigation Equipment are not included in this manual.
These systems are described in separate manuals, available from
the Cessna Service Parts Center.

iii/(iv blank)

SECTION 1
GENERAL DESCRIPTION

Page

TABLE OF CONTENTS
GENERAL DESCRIPTION ..........
Skywagon and Turbo Skywagon 206.....
...
Series .......
........
Description ....
Super Skylane and Turbo Super
Skylane 206-Series ..........

.

1-1

.

1-1
1-1
1-1

1-1. GENERAL DESCRIPTION.
1-2. SKYWAGON AND TURBO SKYWAGON 206-SERIES.

..............
Description
Stationair and Turbo Stationair-Series . .
.......
Description ...........
.
. . .
..
Aircraft Specifications ..
.
Stations ............
......
Torque Values ......

all-metal constant speed propeller. In addition,
Turbo Super Skylane 206-Series engines are turbocharged.
1-6.

1-3. DESCRIPTION. Cessna Skywagon and Turbo
Skywagon 206-Series aircraft, described in this manual, are single-engine, high-wing, strut-braced
monoplanes of all-metal, semimonocoque construction. These aircraft are equipped with a fixed tricycle landing gear employing spring-steel main landing gear struts and a steerable nose gear with an
air/hydraulic fluid shock strut. Wing flaps are electrically-actuated. Both the Skywagon and Turbo
Skywagon 206-Series aircraft are equipped with large
double cargo doors on the right side of the fuselage
and an entrance door on the left side of the cabin.
The pilot's seat only is standard, but provisions are
made for the addition of optional seats to make a sixplace aircraft. Skywagon and Turbo Skywagon 206Series aircraft are powered by a six-cylinder, horizontally opposed, air-cooled, fuel-injection Continental engine, driving an all-metal, constant speed
propeller. In addition, Turbo Skywagon 206-Series
aircraft engines are turbocharged.
1-4. SUPER SKYLANE AND TURBO SUPER SKYLANE 206-SERIES.
1-5. DESCRIPTION. Cessna Super Skylane and
Turbo Super Skylane 206-Series aircraft, described
in this manual, are single-engine, high-wing, strutbraced monoplanes of all-metal, semimonocoque
construction. These aircraft are equipped with a
fixed tricycle landing gear employing spring-steel
main landing gear struts and a steerable nose gear
with an air/hydraulic fluid shock strut. Wing flaps
are electrically-actuated. Both the Super Skylane
and the Turbo Super Skylane 2 0 6-Series aircraft are
equipped with an entrance door'on each side of the
cabin, and a baggage door on the left side of the
fuselage. The seating arrangement of these aircraft
consists of six individual seats. Super Skylane and
Turbo Super Skylane 206-Series aircraft are powered by a six-cylinder, horizontally opposed, aircooled, fuel-injection Continental engine, driving an

1-1
1-1
1-1
1-1
1-1
1-1

STATIONAIR AND TURBO STATIONAIR-SERIES.

1-7. DESCRIPTION. Cessna Stationair and TurboStationair-Series aircraft, described in this manual,
are single-engine, high-wing, strut-braced monoplanes of all-metal, semimonocoque construction.
These aircraft are equipped with a fixed tricycle landing gear employing spring-steel main landing gear
struts and a steerable nose gear with an air/hydraulic
fluid shock strut. Wing flaps are electrically-actuated. Both the Stationair and Turbo Stationair-Series
aircraft are equipped with large double cargo doors
on the right side of the fuselage and an entrance door
on the left side of the cabin. The seating arrangement
of these aircraft consists of six individual seats. Stationair and Turbo Stationair-Series aircraft are powered by a six-cylinder, horizontally opposed, aircooled, fuel-injection Continental engine, driving an
all-metal, constant speed propeller. In addition,
Turbo Stationair engines are turbocharged.
1-8. AIRCRAFT SPECIFICATIONS. Leading particulars of these aircraft, with dimensions based on
gross weight, are given in figure 1-1. If these
dimensions are used for constructing a hangar or
computing clearances, remember that such factors
as nose gear strut inflation, tire pressures, tire
sizes and load distribution may result in some dimensions that are considerably different from those
listed.
1-9. STATIONS. A station diagram is shown in
figure 1-2 to assist in locating equipment when a
written description is inadequate or impractical.
1-10. TORQUE VALUES. A chart of recommended
nut torque values is shown in figure 1-3. These torque values are recommended for all installation procedures contained in this manual, except where other
values are stipulated. They are not to be used for
checking tightness of installed parts during service.
1-1

MODEL P206 AND TP206 SERIES

GROSS WEIGHT .......................
3600 lb
FUEL CAPACITY
Standard Wing (Total) . .
......
. .......
.
65 gal.
Standard Wing (Usable)
......
...
...
....
.
63 gal.
Long-Range Wing ITotal) .
..
............
84 gal.
Long-Range Wing (Usable) .
............
80 gal.
OIL CAPACITY
(Without External Filter) .................
12 qt
(With External Filter) ...................
13 qt
ENGINE MODEL
P206 (Refer to Section 12 for Engine Data) .
........
.
CONTINENTAL 10-520 SERIES
TP206 (Refer to Section 12A for Engine Data)
.........
CONTINENTAL TSIO-520 SERIES
PROPELLER
82" McCAULEY
..
................
Standard (Two Blades) .
Optional (Three Blades) ................
.
80" McCAULEY
MAIN WHEEL TIRES (Standard).
...
. . ..
. ......
6.00 x 6, 6-ply rating
Pressure
. . . . . . . . . . . . . . . . . . . . . . . . 42 psi
MAIN WHEEL TIRES (Optional) . . . . .............
8.00 x 6, 6-ply rating
...
35 psi
. . . . . . . . . ..
Pressure
. . . .... ....
5.00 x 5, 6-ply rating
............
NOSE WHEEL TIRE (Standard) .....
Pressure
. .
.
.
..
. . ... ..
. . . . ..
. 49 psi
6.00 x 6, 6-ply rating
NOSE WHEEL TIRE (Optional) . ...............
Pressure
. . .. . . . .
. . . . . . . . . . . .. . 29 psi
NOSE GEAR STRUT PRESSURE (Strut Extended) .........
80 psi
WHEEL ALIGNMENT
Cam ber . . . . . . . . . . . . . . . . . . . .. .. . . 4 ° ± 1 ° 30'
Toe-In
. . ...
. . .
......
..
. .....
. 0" to .0 6 "
AILERON TRAVEL
21 ± 2 °
....
...............
Up
.......
Down .
........
... ..
. . ...
.
.
14 ° 30' ± 2°
0 ° to 40 ° . + 1 -2 °
..
..
WING FLAP TRAVEL (Electrically-Operated) ..
RUDDER TRAVEL (Measured parallel to water line)
..
24 ° ±1
. . . . . . . . .
. . . . ...
Right. . .
Left ...................
..
24 ° ± 1°
RUDDER TRAVEL (Measured perpendicular to hinge line)
. .
.
27 ° 13' ± I
..
. . . . . . . ..
.. . ....
Right .
..
.
27 ° 13' ±
.
....
. ..
..
..
. ..
Left . . ...
ELEVATOR TRAVEL
.
21 ±1 °
.. . . . . . . . . . . . ..
Up
. . . . . .. . . .
17" ±1 °
. ..
. . . . .......
. ..
Dow n .. . . .
ELEVATOR TRIM TAB TRAVEL
. .. . . .
. . . 25 ° . 41 ° -0O
.
. . .
Up
. . . ..
5 ° +1 -0
..
. .
. ... . . ..............
Dow n .
PRINCIPAL DIMENSIONS
36' 7"
Wing Span (Conventional Wing Tip) .............
35' 10'
...
. .
Wing Span IConical-Camber Wing Tip) ...
.....
13'
.......
Tail Span ...........
.
28' 3"
. .. . . . . . . . . . . . . ......
Length .
.
Fin Height (Maximum with Nose Gear Depressed and
Flashing Beacon Installed on Fin) . ..........
. 9' 7-1/2"
......
. 8' 1-3/4"
.
Track Width . . ...........
..
..
...
Left Side of Firewall
.........
BATTERY LOCATION . ...

Figure 1-1.
1-2

Change 2

Aircraft Specifications (Sheet I of 2)

C^,

112.0

B

100 0

y\

' SUPER SKYLANE 206
100.0

~20601445_^

-j

<12 |
low

^

tf

62 VIEW
-^! *RIGHT 85.
SIDE

X

BEGINNING wITH

8.1
3.8

18.4

NOT USED

OF MODELS WITH
DOORS
<4-p
^
1000CARGO

'

59 70

~1~59.-47

ThRU u?0_

.

44.0 1-2. 8.3
Figure
Wing and Fuselage Reference Stations

65.3

SUPER SKYLANE 206
10.1

112.0

152.2

180.6

THRU U20601444
209.0

ON MODELS WITH
8.1

44.0

Figure 1-2.
1-4

68.3

Wing and Fuselage Reference Stations

MODEL U206 AND TU206 SERIES

GROSS WEIGHT. ...................
.....
.
FUEL CAPACITY
Standard Wing (Total) ....................
.
Standard Wing (Usable) ....................
Long-Range Wing (Total) ..................
..
Long-Range Wing (Usable) .........
..
Standard Wing (Total) ....................
.
Standard Wing (Usable). ............
.
.........
Long-Range Wing (Total). ...............
....
...
Long-Range Wing (Usable) ........
............
.
OIL CAPACITY
(Without External Filter) ................
..
(With External Filter) .
.......
..........
ENGINE MODEL
U206 (Refer to Section 12 for Engine Data) .
...
......
TU206 (Refer to Section 12A for Engine Data)..........
PROPELLER
Standard (Two Blades) .
................
.
Optional (Three Blades) ..................
.
MAIN WHEEL TIRES (Standard).
..........
. . .
Pressure .......................
MAIN WHEEL TIRES (Optional) . ..........
..
Pressure ........................
.
NOSE WHEEL TIRE (Standard) ...............
Pressure .
NOSE WHEEL TIRE (Optional) .....
..
Pressure . . . . . . . . . . . . . . . . . . . . . . . . .
NOSE GEAR STRUT PRESSURE (Strut Extended) ........
WHEEL ALIGNMENT
Camber .
..
...............
....
Toe-In
...........
......
.
.
.
.
AILERON TRAVEL
Up
. . . . . . . . . . . . . . . . . . . . . . . . . . .
Down ..............
WING FLAP TRAVEL (Electrically-Operated)

.

.

.....

. ..

....................
..
..
. . . . ..

65 gal.
63 gal.
84 gal.
80 gal.
61 gal.
59 gal.
80 gal.
76 gal.

When not modified by Cessna
Single-Engine Service Letter
SE75-7 and prior to
U20602127
When modified by Cessna
Single-Engine Service
Letter SE75-7 and beginning with U20602127

12 qt
13 qt
CONTINENTAL 10-520 SERIES
CONTINENTAL TSIO-520 SERIES
82" McCAULEY
80" McCAULEY
6.00 x 6, 6-ply rating
42 psi
8.00 x 6, 6-ply rating
35 psi
5.00 x 5, 6-ply rating
49 psi
6.00 x 6, 6-ply rating
29 psi
80 psi
4' + 1° 30'
0" to .06
21 ° 2 °
14' 30' ± 2'
0

RUDDER TRAVEL (Measured parallel to water line)
Right.
Left ..

3600 lb

to 400, +1

-2

24 ° ± 1'
24 ° ± 1°

.......

RUDDER TRAVEL (Measured perpendicular to hingeline)
Right ..
. . . . . . . . ... . . . . . . . . . .
Left . . . . . . . . . . . . .
.
. . . . . . . . .

...

.
.

27° 13' ± 1
27 ° 13' ± 1°

ELEVATOR TRAVEL
Up
. . . . . . . . . . . . . . . . . . . . . . . . . . .
Dow n . . . . . . . . . . . . . . . . . . . . . . . . . .

ELEVATOR TRIM TAB TRAVEL
Up
. . . . . . . . . . . . . . . . . . ..
Down

.........................

25
.

PRINCIPAL DIMENSIONS
Wing Span (Conventional Wing Tip) ........
.
.
Wing Span (Conical-Camber Wing Tip) .
...........
Tail Span ......................
Length
...................
. . . . . . .
Fin Height (Maximum with Nose Gear Depressed and
Flashing Beacon Installed on Fin) ....
...........
Track Width ..................
.
BATTERY LOCATION (12V) ..................
(24V) Thru 1973) ............
(24V) Beginning with 1974) ........
Figure 1-1.

21 °
17 °

°

5

°

1
1

+1 -0 °
+1 -0

36' 7"
35' 10"
13'
28'

(Add 2" for strobe lights)
g

9' 7-1/2"
8' 1-3/4"
Left side of firewall
Below engine in nose wheel tunnel
Left side of firewall

Aircraft Specifications (Sheet 2 of 2)
Change 3

1-3

RECOMMENDED

NUT TORQUES

THE TORQUE VALUES STATED ARE POUND-INCHES, RELATED
ONLY TO STEEL NUTS ON OIL-FREE CADMIUM PLATED THREADS.
FINE THREAD SERIES
TAP
SIZE
STD
(NOTE 1)
8-36
10-32
1/4-28
5/16-24
3/8-24
7/16-20
1/2-20
9/16-18
5/8-18
3/4-16
7/8-14
1-14
1-1/8-12
1-1/4-12

12-15
20-25
50-70
100-140
160-190
450-500
480-690
800-1000
1100-1300
2300-2500
2500-3000
3700-5500
5000-7000
9000-11000

TENSION

SHEAR

TORQUE

TORQUE

7-9
12-15
30-40
60-85
95-110
270-300
290-410
480-600
660-780
1300-1500
1500-1800
2200-3300
3000-4200
5400-6600

20-28
50-75
100-150
160-260
450-560
480-730
800-1070
1100-1600
2300-3350
2500-4650
3700-6650
5000-10000
9000-16700

ALT
(NOTE 2)

STD
(NOTE 3)

ALT
(NOTE 2)

12-19
30-48
60-106
95-170
270-390
290-500
480-750
660-1060
1300-2200
1500-2900
2200-4400
3000-6300
5400-10000

COARSE THREAD SERIES
(NOTE 5)

(NOTE 4)
8-32
10-24
1/4-20
5/16-18
3/8-16
7/16-14
1/2-13
9/16-12
5/8-11
3/4-10
7/8-9
1-8
1-1/8-8
1-1/4-8

7-9
12-15
25-30
48-55
95-100
140-155
240-290
300-420
420-540
700-950
1300-1800
2200-3000
3300-4000
4000-5000

12-15
20-25
40-50
80-90
160-185
235-255
400-480
500-700
700-900
1150-1600
2200-3000
3700-5000
5500-6500
6500-8000

NOTES
1. Covers AN310, AN315, AN345, AN363, MS20365, MS21042, MS21044, MS21045 and MS21046.
2. When using AN310 or AN320 castellated nuts where alignment between the bolt and cotter pin slots is not
reached using normal torque values, use alternate torque values or replace the nut.
3. Covers AN316, AN320, MS20364 and MS21245.
4. Covers AN363, MS20365, MS21042, MS21043, MS21044, MS21045 and MS21046.
5. Covers AN340.

CAUTION
DO NOT REUSE SELF-LOCKING NUTS.
The above values are recommended for all installation procedures contained in this manual, except where
other values are stipulated. They are not to be used for checking tightness of installed parts during service.

Figure 1-3.

Torque Values
Change 2

1-5/(1-6 blank)

SECTION 2
GROUND HANDLING. SERVICING. CLEANING,

TABLE OF CONTENTS

Page

GROUND HANDLING ...........
Towing . . . . . . . . . . . . . . . .
.
..............
Hoisting
.
............
Jacking
Parking ...............
.
....
......
Tie-Down ....
Flyable Storage ............
.
Returning Aircraft to Service ....
Temporary Storage ..........
Inspection During Storage .....
Returning Aircraft to Service ....
Indefinite Storage ...........
.....
Inspection During Storage
Returning Aircraft to Service . ...
...............
Leveling
...............
SERVICING
Fuel Tanks ....
..........
Fuel Drains .............
...
.
Engine Oil .........
Engine Induction Air Filter .......
Vacuum System Air Filter .......
Battery ............
...

2-1.

LUBRICATION AND INSPECTION

2-1
2-1
2-1
2-1
2-2
2-2
2-2
2-2
2-2
2-4
2-4
2-5
2-5
2-5
2-6
2-6
2-6
2-6
2-6
2-7
2-7
2-7

GROUND HANDLING.

2-2. TOWING. Moving the aircraft by hand is accomplished by using the wing struts and landing gear
struts as push points. A tow bar attached to the nose
gear should be used for steering and maneuvering the
aircraft. When no tow bar is available, press down
at the horizontal stabilizer front spar, adjacent to
the fuselage, to raise the nose wheel off the ground.
With the nose wheel clear of the ground, the aircraft
can be turned by pivoting it about the main wheels.

JCAUTIONI
When towing the aircraft, never turn the nose
wheel more than 35 degrees either side of
center or the nose gear will be damaged. Do
not push on control surfaces or outboard empennage surfaces. When pushing on the tailcone, always apply pressure at a bulkhead to
avoid buckling the skin.
2-3. HOISTING. The aircraft may be lifted with a
hoist of two-ton capacity, either by using hoisting

.....
Tires ....
Nose Gear Strut . . . . . . . . . ..
...
Nose Gear Shimmy Dampener
.
Hydraulic Brake System ...
. . . .....
Oxygen System ....
.............
Face Masks
. ............
CLEANING
.
General Description ....
Upholstery and Interior ..
Plastic Trim ........
Windshield and Windows ....
Aluminum Surfaces .....
.......
Painted Surfaces
.
..
..
.
.....
Engine Compartment
. . . . .....
Propellers .....
...............
Wheels
.
..........
LUBRICATION
General Description ..........
Nose Gear Torque Links ........
.
Tachometer Drive Shaft . . . .
Wheel Bearing Lubrication ...
....
Wing Flap Act ator . .
.
INSPECTION ..............

.

.

2-7
2-8
2-8
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-10
2-10
2-10
2-10
2-10
2-10
2-10
2-10
2-10
2-19

rings (optional equipment) or by using suitable slings.
The front sling should be hooked to the engine lifting
eye, and the aft sling should be positioned around the
fuselage at the first bulkhead forward of the leading
edge of the stabilizer. If the optional hoisting rings
are used, a minimum cable length of 60 inches for
each cable is required to prevent bending of the eyebolt type hoisting rings. If desired, a spreader jig
may be fabricated to apply vertical force to the eyebolts.
2-4. JACKING.
cedures.

Refer to figure 2-2 for jacking proCAUTION|

When using the universal jack point, flexibility
of the gear strut will cause the main wheel to
slide inboard as the wheel is raised, tilting
the jack. The jack must then be lowered for
a second jacking operation. Jacking both
wheels simultaneously with universal jack
points is not recommended.

Change 3

2-1

Figure 2-1.
2-5. PARKING. Parking precautions depend principally on local conditions. As a general precaution,
it is wise to set the parking brake or chock the
wheels, and install the control lock. In severe
weather, and high wind conditions, tie down the aircraft as outlined in paragraph 2-6 if a hangar is not
available.
2-6. TIE-DOWN. When mooring the aircraft in the
open, head into the wind if possible. Secure control
surfaces with the internal control lock and set brakes.

CAUTION
Do not set parking brakes during cold weather
when accumulated moisture may freeze the
brakes or when the brakes are overheated.
a. Tie ropes, cables or chains to the wing tie-down
fittings, located at the upper end of each wing strut.
Secure the opposite ends of ropes, cables or chains
to ground anchors.
b. Secure a tie-down rope (no chains or cables)
to upper trunnion of the nose gear, and secure opposite end of rope to ground anchor.
c. Secure the middle of a rope to the tail tie-down
ring. Pull each end of rope away at a 45-degree
angle and secure to ground anchors at each side of
tail.
d. Secure control lock on pilot control column. If
control lock is not available, tie pilot control wheel
back with front seat belt.
e. These aircraft are equipped with a spring-loaded
steering bungee which affords protection against normal wind gusts. However, if extremely high wind
gusts are anticipated, additional locks may be installed.
2-7. FLYABLE STORAGE. Flyable storage is defined as a maximum of 30 days non-operational stor2-2

Change 2

Typical Tow Bar
age and/or the first 25 hours of intermittent engine
operation.
NOTE
The aircraft is delivered from Cessna with
a Corrosion Preventive Aircraft Engine Oil
(Military Specification MIL-C-6529 Type II
Rust Ban). This engine oil is a blend of aviation grade straight mineral oil and a corrosion preventive compound. This engine oil
should be used for the first 25 hours of engine
operation. Refer to paragraph 2-20 for oil
changes during the first 50 hours of operation.
During the 30 day non-operational storage or the first
25 hours of intermittent engine operation, the propeller shall be rotated through five revolutions every
seventh day, without running the engine. If the aircraft is stored outside, tie it down in accordance
with paragraph 2-6. In addition, the pitot tube,
static air vents, openings in the engine cowling, and
other similar openings shall have protective covers
installed to prevent entry of foreign material. After
30 days, aircraft should be flown for 30 minutes or
ground run-up until oil has reached operating temperature.
2-8. RETURNING AIRCRAFT TO SERVICE. After
flyable storage, returning the aircraft to service is
accomplished by performing a thorough pre-flight inspection. At the end of the first 25 hours of engine
operation, drain engine oil, clean oil screens and
change external oil filter element. Service engine
with correct grade and quantity of oil. Refer to figure 2-4 and paragraph 2-20 for correct grade of
engine oil.
2-9. TEMPORARY STORAGE. Temporary storage
is defined as aircraft in a non-operational status for

ITEM NUMBER

TYPE AND PART NUMBER

II\

Block (Jack point not available)

lx4x4 padded with 1/4 " rubber

Jack

Any short jack of capable capacity

Cessna #SE-767

Universal tail stand (SEE NOTE 1)

O(iJ

REMARKS

Cessna #SE-576 (41-1/2" high)

Universal jack stand (FOR USE WITH ITEM 2)

Cessna #10004-98

Jack point (SEE NOTE 2)

#2-170 Basic jack
#2-109 Leg Extension
#2-70 Slide tube extension

Closed height: 69-1/2 inches: extended
height: 92" Insert slide tube extension
into basic jack)

1. Weighted adjustable stand attaches to tie-down ring.
2.

Cessna #10004-98 jack point may be used to raise only one wheel thru U20602579. Brake line
fairing will prevent jacking aircraft beginning with U20602580 at strut. Do not use brake casting
as a jack point.

3. Items (3), (4). (5) and (6) are available from the Cessna Service Parts Center.
JACKING PROCEDURE
a. Lower aircraft tail so that wing jack can be placed under front spar just outboard of
wing strut.
b. Raise aircraft tail and attach tail stand to tie-down ring. BE SURE the tail stand weighs
enough to keeo the tail down under all conditions and is strong enough to support aircraft
weight.
c. Raise jacks evenly until desired height is reached.
When using the universal jack point, flexibility of the gear strut will cause the main wheel to slide inboard as the wheel is raised, tilting the jack. The jack must be lowered for a second operation. Jacking
both main wheels simultaneously with universal jack points is not recommended.
Figure 2-2. Jacking Details
Change 3

2-3

a maximum of 90 days. The aircraft is constructed
of corrosion resistant alclad aluminum, which will
last indefinitely under normal conditions if kept
clean, however, these alloys are subject to oxidation.
The first indication of corrosion on unpainted surfaces is in the form of white deposits or spots. On
painted surfaces, the paint is discolored or blistered.
Storage in a dry hangar is essential to good preservation and should be procured if possible. Varying
conditions will alter the measures of preservation,
but under normal conditions in a dry hangar, and for
storage periods not to exceed 90 days, the following
methods of treatment are suggested:
a. Fill fuel tanks with correct grade of gasoline.
b. Clean and wax aircraft thoroughly.
c. Clean any oil or grease from tires and coat tires
with a tire preservative. Cover tires to protect
against grease and oil.
d. Either block up fuselage to relieve pressure on
tires or rotate wheels every 30 days to change supporting paints and prevent flat spotting the tires.
e. Lubricate all airframe items and seal or cover
all openings which could allow moisture and/or dust
to enter.
NOTE
The aircraft battery serial number is recorded
in the aircraft equipment list. To assure accurate warranty records, the battery should be
re-installed in the same aircraft from which it
was removed. If the battery is returned to
service in a different aircraft, appropriate
record changes must be made and notification
sent to the Cessna Claims Department.
f. Remove battery and store in a cool dry place;
service the battery periodically and charge as required.

hole of each cylinder with the piston in a down position. Rotate crankshaft as each pair of cylinders
is sprayed.
i. After completing step "h, " rotate crankshaft so
that no piston is at a top position. If the aircraft is
to be stored outside, stop two-bladed propeller so
that blades are as near horizontal as possible to provide maximum clearance with passing aircraft.
j. Again spray each cylinder without moving the
crankshaft to thoroughly cover all interior surfaces
of the cylinder above the piston.
k. Install spark plugs and connect spark plug leads.
1. Apply preservative oil to the engine interior by
spraying approximately two ounces of the preservative
oil through the oil filler tube.
m. Seal all engine openings exposed to the atmosphere using suitable plugs or non-hygroscopic tape.
Attach a red streamer at each point that a plug or
tape is installed.
n. If the aircraft is to be stored outside, perform
the procedures outlined in paragraph 2-6. In addition,
the pitot tube, static source vents, air vents, openings in the engine cowling and other similar openings
should have protective covers installed to prevent
entry of foreign material.
o. Attach a warning placard to the propeller to the
effect that the propeller shall not be moved while the
engine is in storage.
2-10. INSPECTION DURING STORAGE.
a. Inspect airframe for corrosion at least once a
month and remove dust collections as frequently as
possible. Clean and wax as required.
b. Inspect the interior of at least one cylinder
through the spark plug hole for corrosionat least
once a month.
NOTE
Do not move crankshaft when inspecting
interior of cylinder for corrosion.

NOTE
An engine treated in accordance with the following may be considered protected against
normal atmospheric corrosion for a period
not to exceed 90 days.
g. Disconnect spark plug leads and remove upper
and lower spark plugs from each cylinder.
NOTE
The preservative oil must be Lubricating Oil Contact and Volatile, Corrosion Inhibited,
MIL-L-46002. Grade 1 or equivalent. The
following oils are approved for spraying operations by Teledyne Continental Motors,
Nucle Oil 105 - Daubert Chemical Co., 4700
So. Central Ave., Chicago, Illinois, Petratect
VA - Pennsylvania Refining Co., Butler, Pennsylvania, Ferro-Gard 1009G - Ranco Laboratories, Inc., 3617 Brownsville Rd., Pittsburg,
Pennsylvania.
h. Using a portable pressure sprayer, atomize
spray preservative oil through the upper spark plug
2-4

Change 3

c. If at the end of the 90 day period, the aircraft is
to be continued in non-operational storage, again perform the procedural steps "g thru o" of paragraph
2-9.
2-11. RETURNING AIRCRAFT TO SERVICE. After
temporary storage, use the following procedures to
return the aircraft to service.
a. Remove aircraft from blocks and check tires
for proper inflation. Check for proper nose gear strut
inflation.
b. Check battery and install.
c. Check that oil sump has proper grade and quantity
of engine oil.
d. Service induction air filter and remove warning
placard from propeller.
e. Remove materials used to cover openings.
f. Remove, clean, and gap spark plugs.
g. While spark plugs are removed, rotate propeller
several revolutions to clear excess rust preventive
oil from cylinders.
h. Install spark plugs and torque to value specified
in Section 12 or 12A.
i. Check fuel strainer. Remove and clean filter
screen if necessary. Check fuel tanks and fuel lines

for moisture and sediment, drain enough fuel to
eliminate moisture and sediment,
j. Perform a thorough pre-flight inspection, then
start and warm-up engine.
2-12. INDEFINITE STORAGE. Indefinite storage is
defined as aircraft in a non-operational status for an
indefinite period of time. Engines treated in accordance with the following may be considered protected
against normal atmosphere corrosion, provided the
procedures outlined in paragraph 2-13 are performed
at the intervals specified.
at the intervals specified.
a. Operate engine until oil temperature reaches
normal operating range. Drain engine oil sump and
close drain valve or install drain plug.
b. Fill oil sump to normal operating capacity with
corrosion preventive mixture which has been thoroughly mixed and pre-heated to a minimum of 221°F
at the time it is added to the engine.
NOTE
Corrosion-preventive mixture consists of
one part compound MIL-C-6529, Type I,
mixed with three parts new lubricating oil
of the grade recommended for service.
Continental Motors Corporation recommends Cosmoline No. 1223, supplied by
E. F. Houghton & Co., 305 W. LeHigh
Avenue, Philadelphia, Pa. During all
spraying operations corrosionmixture is
pre-heated to 221 ° to 250°F.
c. Immediately after filling the oil sump with corrosion preventive mixture, fy the aircraft for a
period of time not to exceed a maximum of 30 minutes.
d. After flight and with engine operating at 1200
to 1500 rpm and induction air filter removed, spray
corrosion preventive mixture into induction airbox,
at the rate of one-half gallon per minute, until heavy
smoke comes from the exhaust stack, then increase
the spray until engine is stopped.
ICAUTION|
Injecting corrosion-preventive mixture too
fast can cause a hydrostatic lock.
e. Do not rotate propeller after completing step "d."
f. Remove all spark plugs and spray corrosion°
preventive mixture, which has been pre-heated to 221
to 250°F., into all spark plug holes to thoroughly cover interior surfaces of cylinders.
g. Install spark plugs or solid plugs in the lower
spark plug holes and install dehydrator plugs in the
upper spark plug holes. Be sure that dehydrator
plugs are blue in color when installed.
h. Cover spark plug lead terminals with shipping
plugs (AN4060-1) or other suitable covers,
i. With throttle in full open position, place a bag of
desiccant in the induction air intake and seal opening
with moisture resistant paper and tape.
j. Place a bag of desiccant in the exhausts tailpipe
(s) and seal openings with moisture resistant tape.
k. Seal cold air inlet to the heater muff with moisture resistant tape.
1. Seal engine breather tube by inserting a protex

plug in the breather and clamping in place
m. Seal all other engine openings exposed to atmosphere using suitable plugs or non-hygroscopic tape.
N
NOTE
Attach a red streamer to each place plugs
or tape is installed. Either attach red
streamers outside of the sealed area with
tape or to the nside of the sealed area
with safety wire to prevent wicking of
moisture into the sealed area.
n. Drain corrosion-preventive mixture from engine
sump and reinstall drain plug or close drain valve.
NOTE
The corrosion-preventive mixture is harmful to paint and should be wiped from painted
surfaces immediately.
o. Attach a warning placard on the throttle control
knob to the effect that the engine contains no lubricating oil. Placard the propeller to the effect that it
should not be moved while the engine is in storage.
o. Prepare airframe for storage as outlined in
paragraph 2-9 thru step "f."
NOTE
As an alternate method of indefinite storage,
the aircraft may be serviced in accordance
with paragraph 2-9 providing the aircraft is
run-up at maximum intervals of 90 days and
then reserviced per paragraph 2-9.
2-13. INSPECTION DURING STORAGE. Aircraft
in an indefinite storage shall be inspected as follows:
a. Inspect cylinder protex plugs each 7 days.
b. Change protex plugs if their color indicates
an unsafe condition.
c. If the protex plugs have changed color in one
half of the cylinders all desiccant material in the
engine shall be replaced with new material.
d. Every 6 months respray the cylinders interior
with corrosion-preventive mixture.
NOTE
Before spraying, inspect the interior of
one cylinder for corrosion through the
spark plug hole and remove at least one
rocker box cover and inspect the valve
mechanism.
2-14. RETURNING AIRCRAFT TO SERVICE. After
indefinite storage, use the following procedure to return the aircraft to service.
a. Remove aircraft from blocks and check tires for
correct inflation. Check for correct nose gear strut
inflation.
b. Check battery and install.
c. Remove all materials used to seal and cover
openings.
d. Remove warning placards posted at throttle and
2-5

propeller.
e. Remove and clean engine oil screen, then reinstall and safety. On aircraft equipped with an external oil filter, install new filter element,
f. Remove oil sump drain plug or open drain valve
and drain sump. Install or close drain valve and
safety,
NOTE
The corrosion-preventive mixture will mix
with the engine lubricating oil, so flushing
the oil system is not necessary. Draining
the oil sump will remove enough of the corrosion-preventive mixture,
g. Service and install the induction air filter.
h. Remove protex plugs and spark plugs or plugs
installed in spark plug holes and rotate propeller by
hand several revolutions to clear corrosion-preventive mixture from the cylinders.
i. Clean, gap and install spark plugs. Torque plugs
to value specified in Section 12 or 12A.
j. Check fuel strainer. Remove and clean filter
screen. Check fuel tanks and fuel lines for moisture
and sediment, and drain enough fuel to eliminate.
k. Perform a thorough pre-flight inspection, then
start and warm-up engine.
1. Thoroughly clean aircraft and flight test aircraft.
2-15. LEVELING. Reference point for leveling the
aircraft longitudinally is the top centerline of the
tailcone between the rear window and vertical fin.
Corresponding points on front seat rails may be
used to level the aircraft laterally.
2-16.

SERVICING.

2-17. DESCRIPTION. Servicing requirements are
shown in figure 2-4. The following paragraphs
supplement this figure by adding details not included
in the figure.
2-18. FUEL. Fuel cells should be filled immediately after flight to lessen condensation in the cells and
lines. Cell capacities are listed in figure 1-1. The
recommended fuel grade to be used is given in figure
2-4.
2-19. FUEL DRAINS. Drains are located at various
places throughout the fuel system. Refer to Section
13 for locations of the various drains in the system.
The strainer drain valve is an integral part of the
fuel strainer assembly. The strainer drain is equipped with a control which is located adjacent to the
oil dipstick. Access to the control is gained through
the oil dipstick access door. Remove drain plugs
and open drain valves at the intervals specified in
figure 2-4. Also, during daily inspection of the fuel
strainer, if water is found in the strainer. there is a
possibility that the wing cell sumps or fuel lines contain water. Therefore, all fuel plugs should be removed and all water drained from the fuel system.
On aircraft equipped with rubberized fuel cells, a
fuel sampler cup is furnished. To activate drain
valve for fuel sampling, place cup to valve and depress valve with rod protruding from cup. (Refer
2-6

Change 3

to figure 13-5.)
2-20. ENGINE OIL. Check engine lubricating oil
with the dipstick five to ten minutes after the engine
has been stopped. The aircraft should be in as near
a level position as possible when checking the engine
oil so that a true reading is obtained. Engine oil
should be drained while the engine is still hot, and
the nose of the aircraft should be raised slightly for
more positive draining of any sludge which may have
collected in the engine oil sump. Engine oil should
be changed every six months, even though less than
the specified hours have accumulated. Reduce these
intervals for prolonged operations in dusty areas in
cold climates where sludging conditions exist, or
where short flights and long idle periods are encountered, which cause sludging conditions. Always
change oil, clean oil screens and clean and/or change
external filter element whenever oil on the dipstick
appears dirty. Ashless dispersant oil, conforming to
Continental Motors Specification No. MHS-24A, shall
be used in these engines. Multi-viscosity oil may be
used to extend the operating temperature range, improve cold engine starting and lubrication of the engine during the critical warm-up period, thus permitting flight through wider ranges of climate change
without the necessity of changing oil. The multi-viscosity grades are recommended for aircraft engines
subjected to wide variations in ambient air temperatures when cold starting of the engine must be accomplished at temperatures below 30 F.
NOTE
New or newly overhauled engines should be
operated on aviation grade straight mineral
oil until the first oil change. The aircraft is
delivered from Cessna with straight mineral
oil (MIL-C-6529, Type II, RUST BAN.) If
oil must be added during the first 25 hours,
use only aviation grade straight mineral oil
conforming to Specification MIL-6082. After the first 25 hours of operation, drain
engine oil sump and clean both the oil suction
strainer and the oil pressure screen. If an
optional oil filter is installed, change filter
element at this time. Refill sump with
straight mineral oil and use until a total of
50 hours have accumulated or oil consumption has stabilized, then change to ashless
dispersant oil.
When changing engine oil, remove and clean oil
screens, or install a new filter element on aircraft
equipped with an external oil filter. An oil quickdrain valve may be installed. This valve provides
a quick and cleaner method of draining the engine oil.
This valve is installed in the oil drain port of the oil
sump. To drain the oil, proceed as follows:
a. Operate engine until oil temperature is at normal operating temperature.
b. (With Quick-Drain Valve) Attach a hose to the

_

quick-drain valve in oil sump. Push up on quickdrain valve until it locks open, and allow oil to
drain through hose into a container.
c. (Without Quick-Drain Valve) Remove oil drain
plug from engine sump and allow oil to drain into a
container.
d. After oil has drained, close quick-drain valve,
if installed, and remove hose. Install and safety
drain plug.
e. Remove and clean oil screen.
f. Service engine with correct quantity and grade
of engine oil.
~O~~f~en6
oil.
NOTE
Refer to inspection charts for intervals for
changing oil and filter elements. Refer to
figure 2-4 for correct grade of engine oil,
and refer to figure 1-1 for correct capacities.
2-21. ENGINE INDUCTION AIR FILTER. The induction air filter keeps dust and dirt from entering
the induction system. The value of maintaining the
air filter in a good clean condition can never be overstressed. More engine wear is caused through the
use of a dirty or damaged air filter than is generally
believed. The frequency with which the filter should
be removed, inspected and cleaned will be determined
primarily by aircraft operating conditions. A good
general rule, however, is to remove, inspect and
clean the filter at least every 50 hours of engine operating time, and more frequently if warranted by
operating conditions. Some operators prefer to hold
spare induction air filters at their home base of operation so that a clean filter is always readily availfor Under extremely dusty conditions
ableuse.
servicing of the filter i recommended. To
daily
service the induction filter, proceed as follows:
a. Remove filter from aircraft.
NOTE
to filter element
damage
to prevent
Use care
when cleaningreefilter with compressed air.
n c g f r wh c
b. Clean filter by blowing with compressed air (not
over 100 psi) from direction opposite of normal air
100
over
psindicate)
from
direction
oppoiteof
flow. Arrows on filter case indicate direction of
~~normalair fln~ow.
ICAUTION1
Do not use solvent or cleaning fluids to wash
deterhousehold deterwater and
and household
only a
a water
Use only
filter. Use
filter.
gent solution when washing the filter.
r
, te f
.After
as o c gd in sp
c.- Alter >^Tanin as outlned l n irtep "b", the filter
may be washed, if necessary, in a solution of warm
water and a mild household detergent. A cold water
solution may be used.
NOTE
The filter
maybeRemove
assembl
The filter assembly may be cleaned with
compressed air a maximum of 30 times
or it may be washed a maximum of 20

times. A new filter should be installed
after using 500 hours of engine operating
time or one year, whichever should occur
first. However, a new filter should be
installed anytime the existing filter is
damaged. A damaged filter may have
sharp or broken edges in the filtering
panels which would allow unfiltered air
to enter the induction system. Any filter
that appears doubtful, shall have a new
filter installed in its place.
d. After
~gine
washing, rinse filter with clear water
until rinse water draining from filter is clear.
Allow water to drain from filter and dry with compressed air (not over 100 psi).
NOTE
The filtering panels of the filter may become
distorted when wet, but they will return to
their original shape when dry.
e. Be sure airbox is clean, and inspect filter. If
filter is damaged, a new filter should be installed.
I Install filter at entrance to airbox with gasket
on aft face of filter frame and with flow arrows on
filter frame pointed in the correct direction.
AIR FILTER. The vacuum system central air filter keeps dust and dirt
system central air filter keeps dust and dirt
u
nfrom entering the vacuum operated instruments.
spect the filter element every 200 hours of operating
time for damage Change the central air filter element when damaged or at every 500 hours of operating time and whenever the suction gage reading
drops below 4.6 inches of mercury. Also, do not
operate the vacuum system with the filter element
removed or a vacuum line disconnected as particles
of dust or other foreign matter may enter the system
and damage the vacuum operated instruments.
2-23. BATTERY. Battery servicing involves adding
distilled water to maintain the electrolyte even with
the horizontal baffle plate or split ring at the bottom
of the filler holes, checking cable connections, and
neutralizing and cleaning:off any spilled electrolyte
or corrosion. Use bicarbonate of soda (baking soda)
neutralize electrolyte or corrand clean water to neutralize electrolyte or corrogsion. Follow with a thorough flushing with clean water. Do not allow bicarbonate of soda to enter battery. Brighten cable and terminal connection with a
with petroleum jelly before
wire brush, then coat
th battery every 50 hours (or at
wire brushc then
connecting. Check the battery every 50 hours (or at
weather. Add
least every 30 days), oftener in hot weather. Add
only distilled water, not acid or "rejuvenators, " to
maintain electrolyte level in the battery. Inspect the
battery box and clean and remove any evidence of
corrosion.
2-24. TIRES. Maintain tire pressure at the value
specified In figure 1-1. When checking pressure,
examine tires for wear, cuts, bruises and slippage.
oil, grease and mud from tires with soap
and water.

Change 2

2-7

stall valve core in filler valve. Connect valve extension to valve.
h. Infate strut to the pressure specified in figure

~~Y

~ ~~~~~~1-1.

NOTE

/y)^

~

|1~

NOSE GEAR STRUT

Ad.z'i,'"

/

~Check

./^

^-gO._ ^\draulic
·

The nose landing gear shock strut will
normally require only a minimum amount
of service. Maintain the strut extension
pressure as shown in Section 1. Lubricate landing gear as shown in figure 2-5.
the landing gear daily for general
~
cleanliness, security of mounting, and for
hydraulic fluid leakage. Keep machined
surfaces wiped free of dirt and dust, using
a clean lint-free cloth saturated with hy-

fluid (MIL-H-5606) or kerosene.
AU surfaces should be wiped free of excessive hydraulic fluid.

2-26. NOSE GEAR SHIMMYDAMPENER. The
shimmy dampener should be serviced at least every
-. \t i-vV^:^\/100
hours. The dampener must be filled completely
with hydraulic fluid, free of entrapped air with the
· _'sfc<^~,/~~
~compensating piston bottomed in the rod. Check that
piston is completely bottomed as follows:
a. Remove shimmy dampener from the aircraft.
NOTE
b. While holding the shimmy dampener in a vertiValve core remains in strut valve.
cal position with the filler plug pointed upward,
An internal flexible cable, in the
loosen the filler plug.
valve extension, is used to depress
c. Allow the spring to bottom out the floating piston
the valve core in strut valve.
inside the shimmy dampener rod.
When the fluid stops flowing, insert a length of
~~~~d.
-,W~~~~~~~~
stiff wire through the air bleed hole in the setscrew
at the end of the piston rod until it touches the floatFigure 2-3. Strut Filler Valve Extension
ing piston. The depth of insertion should be 3-13/16
NOTE
inches.
'

/_.U

Recommended tire pressures should be maintained. Especially in cold weather, remember
that any drop in temperature of the air inside
a tire causes a corresponding drop in air pressure,
2-25. NOSE GEAR STRUT. The nose gear strut requires periodic checking to ascertain that the strut is
filled with hydraulic fluid and is inflated to the correct air pressure. To fill the nose gear strut with
hydraulic fluid and air, proceed as follows:
a. Weight tail to raise nose wheel off ground.
b. Remove filler valve cap from filler valve or
from lower end of valve extension, and depress valve
core to completely deflate nose strut.
c. Remove valve core from filler valve. It will be
necessary to disconnect filler valve extension from
valve at top of strut.
d. Attach a rubber hose to the filler valve.
e. With other end of rubber hose in a container of
clean hydraulic fluid, compress and extend strut several times. This will draw fluid from container into
the strut, filling strut with hydraulic fluid.
f. After strut has been cycled several times, allow
strut to extend. Holding end of rubber hose above
fluid level in container, slowly compress strut, allowing excess fluid to be drained into container.
g. While strut is compressed, remove hose and in-

2-8

NOTE
If the wire insertion is less than 3-13/16
inches, the floating piston is lodged in the
shaft. If the wire cannot be used to free
the piston, the rod assembly and piston
should be replaced.
Service the shimmy dampener as follows:
a. Remove filler plug from dampener.
b. Move piston completely to opposite end from
filler plug.
c. Fill dampener with clean hydraulic fluid completely full.
d. Reinstall filler plug and safety.
e. Wash dampener in solvent and wipe dry with a
cloth.
f. Reinstall shimmy dampener in aircraft.

NOTE
Keep shimmy dampener, especially the
exposed portions of the dampener piston
shaft, clean to prevent collection of dust
and grit which could cut the seals in the
dampener barrel. Keep machined surfaces wiped free of dirt and dust, using a
clean lint-free cloth saturated with hydraulic fluid (MIL-H-5606) or kerosene.
All surfaces should be wiped free of excessive hydraulic fluid.

2-34. WINDSHIELD AND WINDOWS. These surfaces
should be cleaned carefully with plenty of fresh water
and a mild detergent, using the palm of the hand to
feel and dislodge any caked dirt or mud. A sponge,
soft cloth, or chamois may be used, but only as a
means of carrying water to the plastic. Rinse thoroughly, then dry with a clean moist chamois. Do not
rub the plastic with a dry cloth as this builds up an
electrostatic charge which attracts dust. Oil and
grease may be removed by rubbing lightly with a soft
cloth moistened with Stoddard solvent.
CAUTION

2-27. HYDRAULIC BRAKE SYSTEMS. Check brake
master cylinders and refill with hydraulic fluid as
required every 200 hours. Bleed the brake system of
entrapped air whenever there is a spongy response to
the brake pedals. Refer to Section 5 for filling and
bleeding the brake systems.
2-28.

OXYGEN SYSTEM.

2-29.

FACE MASKS.

Refer to Section 15.

Refer to Section 15.

2-30. CLEANING.
2-31. GENERAL DESCRIPTION. Keeping the aircraft clean is important. Besides maintaining the
trim appearance of the aircraft, cleaning lessens the
possibility of corrosion and makes inspection and
maintenance easier.
2-32. UPHOLSTERY AND INTERIOR. Cleaning
prolongs the life of upholstery fabrics and interior
trim. To clean the interior, proceed as follows:
a. Empty all the ash trays.
b. Brush out or vacuum clean the upholstery and
carpeting to remove dirt.
c. Wipe leather and plastic surfaces with a damp
cloth.
d. Soiled upholstery fabrics and carpet may be
cleaned with a foam-type detergent, used according
to the manufacturer's instructions.
e. Oily spots and stains may be cleaned with household spot removers, used sparingly. Before using
any solvent, read the instructions on the container
and test it on an obscure place in the fabric to be
cleaned. Never saturate the fabric with a volatile
solvent; it may damage the packing and backing material.
f. Scrape off sticky materials with a dull knife,
then spot clean the area.
2-33. PLASTIC TRIM. The instrument panel, plastic trim and control knobs need only be wiped off with
a damp cloth. Oil and grease on the control wheel
and control knobs can be removed with a cloth moistened with Stoddard solvent.
CAUTIONJi
Do not use gasoline, alcohol, benzene, acetone,
carbon tetrachloride, fire extinguisher fluid,
de-icer fluid, lacquer thinner or glass window
cleaning spray. These solvents will soften and
craze the plastic.

Do not use gasoline, alcohol, benzene, acetone,
carbon tetrachloride, fire extinguisher fluid,
de-icer fluid, lacquer thinner or glass window
cleaning spray. These solvents will soften and
craze the plastic.
After washing, the plastic windshield and windows
should be cleaned with an aircraft windshield cleaner.
Apply the cleaner with soft cloths and rub with moderate pressure. Allow the cleaner to dry, then wipe
it off with soft flannel cloths. A thin, even coat of
wax, polished out by hand with soft flannel cloths,
will fill in minor scratches and help prevent further
scratching. Do not use a canvas cover on the windshield or windows unless freezing rain or sleet is
anticipated since the cover may scratch the plastic
surface.
2-35. ALUMINUM SURFACES. The aluminum surfaces require a minimum of care, but should never
be neglected. The aircraft may be washed with clean
water to remove dirt and may be washed with nonalkaline grease solvents to remove oil and/or grease.
Household-type detergent soap powders are effective
cleaners, but should be used cautiously since some
of them are strongly alkaline. Many good aluminum
cleaners, polishes and waxes are available from commercial suppliers of aircraft products.
2-36. PAINTED SURFACES. The painted exterior
surfaces of the aircraft, under normal conditions,
require a minimum of polishing or buffing. Approximately 15 days are required for acrylic paint to cure
completely; in most cases, the curing period will
have been completed prior to delivery of the aircraft.
In the event that polishing or buffing is required within the curing period, it is recommended that the work
be done by an experienced painter. Generally, the
painted surfaces can be kept bright by washing with
water and mild soap, followed by a rinse with water
and drying with cloths or a chamois. Harsh or abrasive soaps or detergents which cause corrosion or
make scratches should never be used. Remove stubborn oil and grease with a cloth moistened with Stoddard solvent. After the curing period, the aircraft
may be waxed with a good automotive wax. A heavier
coating of wax on the leading edges of the wings and
tail and on the engine nose cap will help reduce the
abrasion encountered in these areas.

Change 3

2-9

2-37. ENGINE COMPARTMENT. Cleaning is essential to minimize any danger of fire, and for proper
inspection of engine components. The engine and
engine compartment may be washed down with a
suitable solvent, such as Stoddard solvent or equivalent, then dried thoroughly.
CAUTION
Particular care should be given to electrical
equipment before cleaning. Solvent should
not be allowed to enter magnetos, starters,
alternators, voltage regulators, and the like.
Hence, these components should be protected
before saturating the engine with solvent.
Any oil, fuel, and air openings on the engine
and accessories should be covered before
washing the engine with solvent. Caustic
cleaning solutions should be used cautiously
and should always be properly neutralized
after their use.
2-38. PROPELLER. The propeller should be wiped
occasionally with an oily cloth to remove grass and
bug stains. In salt water areas, this will assist in
corrosion-proofing the propeller.
2-39. WHEELS. The wheels should be washed
periodically and examined for corrosion, chipped
paint and cracks or dents in the wheel castings.
Sand smooth, prime and repaint minor defects.
Cracked wheel halves shall be replaced.
2-40.

LUBRICATION.

2-41. GENERAL DESCRIPTION. Lubrication requirements are outlined in figure 2-5. Before adding lubricant to a fitting, wipe the fitting free of
dirt. Lubricate until grease appears around part
being lubricated and wipe excess grease from parts.
The following paragraphs supplement figure 2-5 by
adding details not shown in the figure.
2-42. NOSE GEAR TORQUE LINKS. Lubricate
torque links every 50 hours. When operating in
dusty conditions, more frequent lubrication is recommended.
2-43. TACHOMETER DRIVE SHAFT.
tion 16 for lubrication instructions.

Refer to Sec-

2-44. WHEEL BEARING LUBRICATION. Clean and
repack wheel bearings at the first 100-hour inspection
and at each 500-hour inspection thereafter. If more
than the usual number of take-off and landings are
made, extensive taxiing is required or the aircraft
is operated in dusty areas or under seacoast conditions, clean and lubricate wheel bearings at each
100-hour inspection.
2-45. WING FLAP ACTUATOR
a. On aircraft prior to P20600648 and U20601673
which have not been modified by Service Kit SK15037, proceed as follows:
1. At each 100 hour inspection, inspect wing
flap actuator jack screw and ball retainer assembly
for lubrication, and lubricate if required. Also,
2-10

remove, clean and lubricate jack screw whenever
actuator slippage is experienced. If lubrication is
required, proceed as follows:
a. Gain access to actuator by removing
appropriate inspection plates on lower surface of
wing.
b. Expose jack screw by operating flaps to
full-down position.
c. Wipe a small amount of lubricant from
jack screw with a rag and examine for condition.
Lubricant should not be dirty, sticky, gummy or
frothy in appearance.
d. Inspect wiped area on jack screw for
presence of hard scale deposit. Previous wiping
action, will have exposed bare metal if no deposit
is present.
e. If any of the preceding conditions exist,
clean and relubricate jack screw as outlined in steps
"f" thru "n".
f. Remove actuator from aircraft in accordance with procedures outlined in Section 7.
g. Remove all existing lubricant from jack
screw and torque tube by running the nut assembly
to the end of the jack screw away from the gearbox,
and soaking the nut assembly and jack screw in Stoddard solvent.
NOTE
Care must be taken to prevent solvent from
entering gearbox. The gearbox lubricant is
not affected and should not be disturbed.
h. After soaking, clean entire length of
jack screw with a wire brush, rinse with solvent
and dry with compressed air.
NOTE
Do not disassemble nut and ball retainer
assembly.
i. Relubricate jack screw with MIL-G21164 (Molybdenum Disulfide Grease) as outlined in
steps "j" thru "m".
j. Rotate nut down screw toward the motor.
k. Coat screw and thread end of nut with
grease and run nut to full extension.
1. Repeat the process and pack lubricant in
the cavity between the nut and ball retainer at the
threaded end of the nut.
m. Repeat the process and work nut back and
forth several times.
n. Remove excess grease.
o. Reinstall actuator in aircraft in accordance with instructions outlined in Section 7.
b. On aircraft prior to Serials P20600648 and U206601673 which have been modified by Service Kit
SK150-37 proceed as follows:
1. At each 100-hour inspection, expose jack
screw by operating flaps to full-down position, and
inspect wing flap actuator jack screw for proper
lubrication. If lubrication is required, proceed as
follows:
a. Clean jack screw with solvent rag, if
necessary, and dry with compressed air.
b. Relubricate jack screw with MIL-G-

21164 (Molybdenum Disulfide Grease) as required.
c. On aircraft beginning with Serial U20601673,
clean and lubricate wing flap actuator jack screw

It is not necessary to remove actuator from

each 100 hours as follows:

aircraft to clean or lubricate threads.

1. Expose jack screw by operating flaps to fulldown position.
2. Clean jack screw threads with solvent rag
and dry with compressed air.

NOTE

3. With oil can, apply light coat of No. 10
weight, non-detergent oil to threads of jack screw.

SHOP NOTES:

2-11

HYDRAULIC FLUID:
SPEC. NO. MIL-H-5606
OXYGEN:
SPEC. NO. MIL-0-27210
RECOMMENDED FUEL:
ENGINE MODEL 10-520 Series CONTINENTAL
FUEL:

Compliance with conditions stated in Continental aircraft engine Service Bulletins
M74-6 and M75-2 and supplements or revisions thereto, are recommended when
using alternate fuel.
1. MINIMUM: 100/130 Aviation Grade
2. ALTERNATE:
a. 115/145 Aviation Grade (with lead content limited to a maximum of 4.6 cc Tetraethyl
lead per gallon.)

Figure 2-4.
2-12

Change 3

Servicing (Sheet 1 of 3)

RECOMMENDED ENGINE OIL:
ENGINE MODEL IO-520-Series CONTINENTAL
AVIATION GRADE:
SAE 50
SAE 30

40°F
40°F

Aviation grade ashless dispersant oil, conforming to Continental Motors Specification MHS-24
and all revisions and supplements thereto, must be used except as noted in paragraph 2-20.
Refer to Continental aircraft Engine Service Bulletin M75-2 and any superseding bulletins,
revisions or supplements thereto, for further recommendations.

DAILY

3

FUEL CELLS:
Service after each flight.

6

FUEL CELL SUMP DRAINS:
Drain off any water and sediment before first flight of the day.

18

FUEL STRAINER:
Drain off any water and sediment before first flight of the day.

15

OIL DIPSTICK:

8
7
4

Keep full to retard condensation.

Check on preflight. Add oil as necessary.
filler cap is tight and oil filler is secure.

Refer to paragraph 2-18 for details.

Refer to paragraph 2-20 for details.

Check that

PITOT AND STATIC PORTS:
Check for obstructions before first flight of the day.
OXYGEN CYLINDERS:
Check for anticipated requirements before each flight.
INDUCTION AIR FILTER:
Inspect and service under dusty conditions.

Refer to Section 15 for details.

Refer to paragraph 2-21 for details.

FIRST 25 HOURS

19

ENGINE OIL SYSTEM:
Refill with straight mineral oil, non-detergent, and use until a total of 50 hours have accumulated or oil consumption has stabilized, then change to ashless dispersant oil.
50 HOURS

4

INDUCTION AIR FILTER:

Clean per paragraph 2-21.

Replace as required.

13

BATTERY:

19

ENGINE OIL SYSTEM:

16

SHIMMY DAMPENER:
Check fluid level and refill as required in accordance with paragraph 2-26.

9

TIRES:
Maintain correct tire inflation as listed in figure 1-1.

Check electrolyte level and clean battery compartment each 50 hours or 30 days.
Change oil each 50 hours if engine is NOT equipped with external filter; if equipped with
external oil filter, change filter element each 50 hours and oil at least at each 100 hours,
or every 6 months.

Figure 2-4.

Refer to paragraph 2-24.

Servicing (Sheet 2 of 3)
Change 3

2-13

50 HOURS (Cont.)
17

20

NOSE GEAR SHOCK STRUT:
Keep strut filled and inflated to correct pressure.
SPARK PLUGS:
Remove, clean and re-gap all spark plugs.

Refer to paragraph 2-25.

Refer to Section 12 or 12A.

100 HOURS
VACUUM SYSTEM OIL SEPARATOR:
Remove, flush with solvent, and dry with compressed air.

1

2

FUEL/AIR CONTROL UNIT SCREEN:
Remove and clean screen.

5

VACUUM RELIEF VALVE FILTER SCREEN:
Remove, flush with solvent and dry with compressed air.

18

22

FUEL STRAINER:

Disassemble and clean strainer bowl and screen.
ALTERNATOR SUPPORT BRACKET:
Check alternator support bracket for security and cracking.
Also refer to Service Letter SE71-42.

200

HOURS

6

FUEL CELL SUMP DRAINS:
Drain off any water or sediment.

10

FUEL RESERVOIR TANK AND/OR SELECTOR VALVE DRAINS:
Remove plugs and drain off any water and sediment. Reinstall and resafety plugs.

12

BRAKE MASTER CYLINDERS:
Check fluid level and fill as required with hydraulic fluid.

0
11

VACUUM SYSTEM CENTRAL AIR FILTER:
Replace every 500 hours.
/

14

HOURS

5500

AS REQUIRED

GROUND SERVICE RECEPTACLE:
Connect to 12-volt, or 24-volt if aircraft is equipped with a 24-volt battery, DC,
negative-ground power unit for cold weather starting and lengthy ground maintainance
of the aircraft electrical equipment with the exception of electronic equipment.
Master switch should be turned on before connecting a generator type or battery
type external power source.
NOTE
The ground power receptacle circuit incorporates a polarity reversal
protection. Power from the external power source will flow only if the
ground service plug is connected correctly to the aircraft.
Figure 2-4.

2-14

Change 3

Servicing (Sheet 3 'f 3)

FREQUENCY (HOURS)
FREQUENCY~ (HOURS)

50^eALSO

0C

A
_,

REFER TO

I PARAGRAPH 2-42

)

TORQUE LINKS

NEEDLE BEARING
Go (STEERING COLLAR)

nil
NOSE GEAR

NOSE WHEEL
BEARINGS

Figure 2-5.

BEARIN SM
II
|
^iy--~REFER
MAIN WHEEL
BEARINGS

AIN GEAR
TO
PARAGRAPH 2-44

Lubrication (Sheet 1 of 4)
Change 1

2-15

NEEDLE BEARINGS

.

-1 OILITE BEARINGS
06

AI LERON DRIVE
PULLEYS

ALL PIANO
HINGES
BATTERY TERMINALS

OILITE BEARINGS

THREADS

ALSO REFER TO INSPECTION
IN THIS SECTION AND
TO SECTION 9 OF THIS MANUAL.

NOSE GEAR
-CHART
BUNGEE
GREASE SPARINGLY

ELEVATOR TRIM
TAB ACTUATOR

Figure 2-5.
2-16

Change 1

Lubrication (Sheet 2 of 4)

ELECTRIC FLAP
DRIVE MECHANISM
AILERON BELLCRANKS

NEEDLE
BEARINGS

CONTROL COLUMN

FLAP BELLCRANKS
AND DRIVE PULLEYS

NEEDLE BEARINGS
WING STRUT-ATTACH
(LOWER) BOLT & HOLE*
*UPON INSTALLATION

Figure 2-5.

Lubrication (Sheet 3 of 4)
Change 1

2-17

RUDDER BARS AND PEDALS
i *r

PARKING BRAKE
HANDLE SHAFT

r BEARING BLOCK
*»
HALVES
06^HALVES

06
OILITE BEARINGS
(RUDDER BAR ENDS)

s

ALL LINKAGE
POINT PIVOTS 0C

/
-~

>XI,

|1|
^

\

Lubricate between inside face of
on shaft and drum

~washer

Lubricate between
washer and drum

BEGINNING
WITH 20601701

:

THRU P20600648
& U20601700/

/

'

Lubricate between inside face of
washer on shaft and drum
Lubricate shaft and small gear
with clutch in open position

\-NOTE
v! r .

i••,^<~.
~

§ ' ~\
|~

ELECTRIC TRIM
ASSEMBLY

Drum groove and cable must
be free of grease and oil
NOTES

Sealed bearings require no lubrication.
McCauley propellers are lubricated at overhaul and require no other lubrication.
Do not lubricate roller chains or cables except under seacoast conditions.
dry cloth.

Wipe with a clean,

Lubricate unsealed pulley bearings, rod ends, Oilite bearings, pivot and hinge points, and any
other friction point obviously needing lubrication, with general purpose oil every 1000 hours or
oftener if required.
Paraffin wax rubbed on seat rails will ease sliding the seats fore and aft.
Lubricate door latching mechanism with MIL-G-81322A or equivalent lubricant,
applied sparingly to friction points, every 1000 hours or oftener if binding occurs.
No lubrication is recommended on the rotary clutch.
Figure 2-5.
2-18

Lubrication (Sheet 4 of 4)

I

INSPECTION REQUIREMENTS.

As required by Federal Aviation Regulations, all civil aircraft of U.S. registry must undergo a
COMPLETE INSPECTION (ANNUAL) each twelve calendar months. In addition to the required
ANNUAL inspection, aircraft operated commercially (for hire) must also have a COMPLETE
AIRCRAFT INSPECTION every 100 hours of operation.
In lieu of the above requirements, an aircraft may be inspected in accordance with a
progressive inspection schedule, which allows the work load to be divided into smaller
operations that can be accomplished in shorter time periods.
Therefore, the Cessna Aircraft Company recommends PROGRESSIVE CARE for aircraft that
are being flown 200 hours or more per year, and the 100 HOUR inspection for all other aircraft.
11

INSPECTION CHARTS.

The following charts show the recommended intervals at which items are to be inspected.
As shown in the charts, there are items to be checked each 50 hours, each 100 hours, each
200 hours, and also Special Inspection items which require servicing or inspection at
intervals other than 50, 100 or 200 hours.

III

a.

When conducting an inspection at 50 hours, all items marked under EACH 50 HOURS would be
inspected, serviced or otherwise accomplished as necessary to insure continuous
airworthiness.

b.

At each 100 hours, the 50 hour items would be accomplished in addition to the items
marked under EACH 100 HOURS as necessary to insure continuous airworthiness.

c.

An inspection conducted at 200 hour intervals would likewise include the 50 hour
items and 100 hour items in addition to those at EACH 200 HOURS.

d.

The numbers appearing in the SPECIAL INSPECTION ITEMS column refer to data listed
at the end of the inspection charts. These items should be checked at each inspection
interval to insure that applicable servicing and inspection requirements are accomplished
at the specified intervals.

e.

A COMPLETE AIRCRAFT INSPECTION includes all 50, 100 and 200 hour items plus those
Special Inspection Items which are due at the time of the inspection.

INSPECTION PROGRAM SELECTION.

AS A GUIDE FOR SELECTING THE INSPECTION PROGRAM THAT BEST
SUITS THE OPERATION OF THE AIRCRAFT, THE FOLLOWING IS
PROVIDED.
1.

IF THE AIRCRAFT IS FLOWN LESS THAN 200 HOURS ANNUALLY.
. IF FLOWN FOR HIRE
An aircraft operating in this category must have a COMPLETE AIRCRAFT INSPECTION
each 100 hours and each 12 calendar months of operation. A COMPLETE AIRCRAFT
INSPECTION consists of all 50, 100, 200 and Special Inspection Items shown in the inspection charts as defined in paragraph i above.
b. IF NOT FLOWN FOR HIRE
An aircraft operating in this category must have a COMPLETE AIRCRAFT INSPECTION each
12 calendar months (ANNUAL). A COMPLETE AIRCRAFT INSPECTION consists of all 50,
100, 200 and Special Inspection Items shown in the inspection charts as defined in paragraph n
above. In addition, it is recommended that between annual inspections, all items be inspected
at the intervals specified in the inspection charts.

Change 3

2-19

2.

IF THE AIRCRAFT IS FLOWN MORE THAN

200 HOURS ANNUALLY.

Whether flown for hire or not, it is recommended that aircraft operating in this category
be placed on the CESSNA PROGRESSIVE CARE PROGRAM. However, if not placed on
Progressive Care, the inspection requirements for aircraft in this category are the
same as those defined under paragraph III 1. (a) and (b).
Cessna Progressive Care may be utilized as a total concept program which
insures that the inspection intervals in the inspection charts are not exceeded.
Manuals and forms which are required for conducting Progressive Care inspections are available from the Cessna Service Parts Center.
IV

INSPECTION GUIDE LINES.
(a) MOVABLE PARTS for: lubrication, servicing, security of attachment, binding, excessive wear,
safetying, proper operation, proper adjustment, correct travel, cracked fittings, security of
hinges, defective bearings, cleanliness, corrosion, deformation, sealing and tension.
(b)

FLUID LINES AND HOSES for: leaks, cracks, dents, kinks, chafing, proper radius, security,
corrosion, deterioration, obstruction and foreign matter.

(c)

METAL PARTS for: security of attachment, cracks, metal distortion, broken spotwelds,
corrosion, condition of paint and any other apparent damage.

(d) WIRING for: security, chafing, burning, defective insulation, loose or broken terminals,
heat deterioration and corroded terminals.
(e)

BOLTS IN CRITICAL AREAS for: correct torque in accordance with torque values given in the
chart in Section 1, when installed or when visual inspection indicates the need for a
torque check.
NOTE
Torque values listed in Section 1 are derived from oil-free cadmium-plated threads,
and are recommended for all installation procedures contained in this book except
where other values are stipulated. They are not to be used for checking tightness of
installed parts during service.

(f)

FILTERS, SCREENS & FLUIDS for: cleanliness, contamination and/or replacement at specified
intervals.

(g)

AIRCRAFT FILE.
Miscellaneous data, information and licenses are a part of the aircraft file. Check that
the following documents are up-to-date and in accordance with current Federal
Aviation Regulations. Most of the items listed are required by the United States
Federal Aviation Regulations. Since the regulations of other nations may require
other documents and data, owners of exported aircraft should check with their
own aviation officials to determine their individual requirements.
To be displayed in the aircraft at all times:
1. Aircraft Airworthiness Certificate (FAA Form 8100-2).
Aircraft Registration Certificate (FAA Form 8050-3).
2.
3. Aircraft Radio Station License, if transmitter is installed (FCC Form 556).
To be carried in the aircraft at all times:
1. Weight and Balance, and associated papers (Latest copy of the Repair and Alteration
Form, FAA Form 337, if applicable).
2.
Aircraft Equipment List.
To be made available upon request:
1. Aircraft Log Book and Engine Log Book.

2-20

Change 1

(h)

ENGINE RUN-UP.
Before beginning the step-by-step inspection, start, run up and shut down the engine in
accordance with instructions in the Owner's Manual. During the run-up, observe the
following, making note of any discrepancies or abnormalities:
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.

Engine temperatures and pressures.
Static rpm. (Also refer to Section 12 or 12A of this Manual.)
Magneto drop. (Also refer to Section 12 or 12A of this Manual).
Engine response to changes in power.
Any unusual engine noises.
Fuel selector and/or shut-off valve; operate engine(s) on each tank (or cell) position
and OFF position long enough to ensure shut-off and/or selector valve functions
properly.
Idling speed and mixture; proper idle cut-off.
Alternator and ammeter.
Suction gage.
Fuel flow indicator.

After the inspection has been completed, an engine run-up should again be performed to determine
that any discrepancies or abnormalities have been corrected.

SHOP NOTES:

Change 1

2-21

SPECIAL INSPECTION ITEM
IMPORTANT

EACH 200 HOURS
EACH 100 HOURS
EACH 100 HOURS
EACH 50 HOURS

READ ALL INSPECTION REQUIREMENTS PARAGRAPHS PRIOR TO
USING THESE CHARTS.
PROPELLER

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1.

Spinner

2.

Spinner bulkhead . ...................

3.

Blades

4.

Bolts and Nuts

5.

Hub

6.

Governor and control

. . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...
............................

.

ENGINE COMPARTMENT
Check for evidence of oil and fuel leaks, then clean entire engine and
compartment, if needed, prior to inspection.
1.

Engine oil screen, filler cap, dipstick, drain plug and external filter element

2.

Oil Cooler

3.

Induction air filter

4.

Induction airbox, air valves, doors and controls

5.

Cold and hot air hoses . ..

6.

Engine baffles

7.

Cylinders, rocker box covers and push rod housings

8.

Crankcase, oil sump, accessory section and front crankshaft seal

9.

Hoses, metal lines and fittings

*

. ..

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...
.....

.

.

........................

. . ..

. . ..

2
.

................
. . ..

. .
..

..................

..

..

. . . ..

. ..

.

..............
..

..

3

... ....

.....................

. . ..

.

......

...

...

. ...

. . ..

Intake and exhaust systems

11.

Ignition harness ..........................

12.

Spark plugs ..........................

13.

Compression check

14.

Crankcase and vacuum system breather lines

15.

Electrical wiring

16.

Vacuum pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

17.

Vacuum relief valve filter

18.

Engine controls and linkage

19.

Engine shock mounts, mount structure and ground straps

Change 1

4

. .

. . ..

10.

2-22

I

.
*

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
.........

...................

.

*

.........

...................

...

........................
........

·

5
6

SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
20. Cabin heat valves, doors and controls .............................................
21. Starter, solenoid and electrical connections ........................................
22. Starter brushes, brush leads and commutator ...................

...................

23. Alternator and electrical connections ..............................................

7

24. Alternator brushes, brush leads, and commutator or slip ring .........................
25. Voltage regulator mounting and electrical leads ....................................
26. Magnetos (external) and electrical connections ............................
27. Magneto timing ..............

·
.........

...........................................

8

28. Fuel-air (metering) control unit ...................................................
29. Firewall.......................................................................
30. Fuel injection system ...........................................................

·

31. Engine cowl flaps and controls ...................................................

a

32. Engine cowling ........................................

.......................

*

33. Turbocharger..................................................................

*

9

34. All oil lines to turbocharger, waste gate and controller ................................
35. Waste gate, actuator and controller ...............................................

0

36. Turbocharger pressurized vent lines to fuel pump, discharge nozzles
and fuel flow gage ..................................

.................

...........

37. Turbocharger mounting brackets and linkage ......................................
38. Alternator support bracket for security (Refer to Service Letter SE71-42) ..............
FUEL SYSTEM
1. Fuel strainer, drain valve and control, cell vents, caps and placards ...................

·

2. Fuel strainer screen and bowl ....................................................

·

3. Fuel injector screen ............................................................

·

4. Fuel reservoirs .................................................................

·

5. Drain fuel and check cell interior, attachment and outlet screens ......................

5

6. Fuel cells and sump drains ......................................................

·

7. Fuel selector valve and placards .................................................
8. Auxiliary fuel pump ...................................................

D2007-3-13 Temporary Revision Number 5 - Jan 6/2003
0 Cessna Aircraft Company

0

..... ...

Change 1

·

2-23

SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
9. Engine-driven fuel pump ........................................................

·

10. Fuel quantity indicators and transmitters ...........................................

|

11. Vapor return line and check valve ................................................

·

12. Turbocharger vent system .......................................................

·

13. Engine primer .................................................................

·

14. Perform a fuel quantity indicating system operational test. Refer to Section 16
for detailed accomplishment instructions .........................................

17

15. Fuel injection nozzles ..........................................................

19

LANDING GEAR
1. Brake fluid, lines and hoses, linings, disc, brake assemblies and master cylinders .......
2. Main gear wheels ..................................

.................

3. Wheel bearings .................

..............................................

4. Main gear springs ................

.................

0

........
10

..........................

.

5. Tires .........................................................................
6. Torque link lubrication ..........................................................
7. Parking brake system .................

.................

.................

.....

8. Nose gear strut and shimmy dampener (service as required) .........................
9. Nose gear wheel ..................................

.................

10. Nose gear fork ....................................................
11. Nose gear steering system .................

.........
...........

................................

12. Parking brake and toe brakes operational test ......................................
AIRFRAME

1. Aircraft exterior ................................................... .............
2. Aircraft structure .........................................................
3. Windows, windshield, doors and seals ........................................
4. Seat stops, seat rails, upholstery, structure and mounting ............................
5. Control column bearings, pulleys, cables, chains and turnbuckles .....................
6. Seat belts and shoulder harnesses ...............................................
7. Control lock, control wheel and control mechanism .................................
8. Instruments and markings .......................................................
9. Gyros central air filter ..........................................................

2-24

11

D2007-3-13 Temporary Revision Number 6 - Apr 5/2004
© Cessna Aircraft Company

SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS

181

10. Magnetic compass compensation ................................................
11. Instrument wiring and plumbing ..........

12.
13.
14.
15.
16.
17.
18.

........................................

Instrument panel, shock mounts, ground straps, cover, decals and labeling.............
Defrosting, heating and ventilating systems and controls .............................
Cabin upholstery, trim, sun visors and ash trays ....................................
Area beneath floor, lines, hose, wires and control cables .............................
Lights, switches, circuit breakers, fuses, and spare fuses ............................
Exterior lights ..................................................................
Pitot and static systems .........................................................

*
*
*

19. Stall warning unit and pitot heater .................................................
20. Radios, radio controls, avionics and flight instruments ...............................
21. Antennas and cables ...........................................................

*

22. Battery, battery box and battery cables ............................................

12

23. Battery electrolyte ..............................................................

*

24. Emergency locator transmitter ...................................................

13

25. Oxygen system ................................................................
26. Oxygen supply, masks and hose .................................................

*

14

27. Inspect all fluid carrying lines and hoses in the cabin and wing areas
for leaks, damage, abrasion, and corrosion ........................................
CONTROL SYSTEMS
In addition to the items listed below, always check for correct direction of movement,
correct travel and correct cable tension.
1. Cables, terminals, pulleys, pulley brackets, cable guards, turnbuckles and fairleads......

2. Chains, terminals, sprockets and chain guards .....................................
3. Trim control wheels, indicators, actuator and bungee ................................
4. Travel stops ...................................................................
*

5. Decals and labeling.............................................................
6. Flap control switch, flap rollers and flap position indicator ............................
7. Flap motor, transmission, limit switches, structure, linkage, bellcranks etc...........
8. Flap actuator jackscrew threads ..................................................

.
.

15

9. Elevators, trim tab, hinges and push-pull tube ......................................
10. Elevator trim tab actuator lubrication and tab free-play inspection .....................

Temporary Revision Number 5
6 January 2003

©2003 CESSNA AIRCRAFT COMPANY

16

2-25

SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
11.

Rudder pedal assemblies and linkage .............................................

12. Extenal skins of control surfaces and tabs.........................................
13. Internal structure of control surfaces ........................................

......

14. Balance weight attachment ......................................................
SPECIAL INSPECTION ITEMS
1.

First 25 hours: use mineral oil confirming witn MIL-C-6529 Type II forthe first 25 nours of operation or
until oil consumption has stabilized, or six months, whichever occurs first. If oil consumption has not
stabilized in this time, drain and replenish the oil and replace the oil filter. After the oil consumption has
stabilized, change to an ashless dispersant oil, refer to Teledyne Continental Service Information Letter
SIL99-2, or latest revision for a current listing of lubricants authorized by TCM. Change oil each 25 hours
if engine is NOT equipped with external oil filter; if equipped with an external oil filter, change oil filter
element and oil at each 50 hours of operation or every six months, whichever occurs first. Refer to the
latest edition of the TCM engine operator/maintenance manual for the latest oil change intervals and
inspection procedures.

2.

Clean filter per paragraph 2-21. Replace as required.

3.

Replace engine compartment hoses per the following schedule:
A. Cessna Installed Flexible Fluid Carrying Rubber Hoses; replace every 5 years or at engine overhaul,
whichever occurs first.
B.

Cessna Installed Flexible Fluid Carrying Teflon Hoses, replace every 10 years or at engine overhaul,
whichever occurs first.

C. TCM Installed Engine Compartment Flexible Fluid Carrying Hoses, refer to Teledyne Continental
Service Bulletin SB97-6 or latest revision for hose replacement intervals.
4.

General inspection every 50 hours. Refer to Section 12 and 12A for 100 hour inspection.

5. Each 1000 hours, or at engine overhaul, whichever occurs first.
6.

Each 50 hours for general condition and freedom of movement. These controls are not repairable,
replace throttle, propeller, and mixture controls at each engine overhaul.

7.

Each 500 hours.

8.

Internal timing and magneto-to-engine timing are described in detail in Section 12.

9.

Remove insulation blanket or heat shields and inspect for burnt area, bulges or cracks. Remove tailpipe
and ducting; inspect turbine for coking, carbonization, oil deposits and turbine impeller for damage.

10.

First 100 hours and each 500 hours thereafter. More often if operated under prevailing wet of dusty
conditions.

11.

Replace each 500 hours.

12. Check electrolyte level and clean battery compartment each 50 hours or 30 days, whichever occurs first.

2-26

©2003 CESSNA AIRCRAFT COMPANY

Temporary Revision Number 5
6 January 2003

13.

Refer to Section 17;

14.

Inspect masks, hose and fittings for condition, routing and support. Test, operate, and check for leaks.

15.

Refer to paragraph 2-45 for detailed instructions for various serial ranges.

16.

Replacement or overhaul of the actuator is required each 1000 hours and/or 3 years, whichever comes
first. Refer to figure 2-5 for grease specifications.
NOTE: Refer to Section 9 of this service manual and Cessna Single Engine Service Letter SE73-25, or
latest revision, for free-play limits, inspection, replacement and/or repair information.

17.

Fuel quantity indicating system operational test is required every 12 months. Refer to Section 16 for
detailed accomplishment instructions.

18. Every 2 years, or anytime components are added or removed which have the potential to affect the
magnetic accuracy and/or variation of the compass calibration, or anytime the accuracy of the compass
is in question. If required, refer to AC 43.13-1 B for compass swing procedures.
19.

2-46.

At the first 100-hour inspection on new, rebuilt or overhauled engines, remove and clean the fuel injection
nozzles. Thereafter, the fuel injection nozzles must be cleaned at 300-hour intervals or more frequently if
fuel stains are found.
COMPONENT TIME LIMITS
1. General
A.

Most components listed throughout Section 2 should be inspected as detailed elsewhere in
this section and repaired, overhauled or replaced as required. Some components, however,
have a time or life limit, and must be overhauled or replaced on or before the specified time
limit.
NOTE:

The terms overhaul and replacement as used within this section are defined as
follows:
Overhaul - Item may be overhauled as defined in FAR 43.2 or it can be replaced.
Replacement - Item must be replaced with a new item or a serviceable item that is
within its time and serviceable life limits or has been rebuilt as defined in FAR 43.2.

B. This section provides a list of items which must be overhauled or replaced at specific time
limits. Table 1 lists those items which Cessna has mandated must be overhauled or replaced
at specific time limits. Table 2 lists component time limits which have been established by a
supplier to Cessna for the supplier's product.
C.

2.

In addition to these time limits, the components listed herein are also inspected at regular time
intervals set forth in the Inspection Charts, and may require overhaul/replacement before the
time limit is reached, based on service usage and inspection results.

Cessna-Established Replacement Time Limits.
A. The following component time limits have been established by Cessna Aircraft Company.
Table 1: Cessna-Established Replacement Time Limits
COMPONENT

REPLACEMENT
TIME

Restraint Assembly Pilot, Copilot,
and Passenger Seats

10 years

D2007-3-13 Temporary Revision Number 6 - Apr 5/2004
© Cessna Aircraft Company

OVERHAUL
NO
2-27

2-28

OVERHAUL

COMPONENT

REPLACEMENT
TIME

Trim Tab Actuator

1,000 hours or 3 years,
whichever occurs first

YES

Vacuum System Filter

500 hours

NO

Vacuum System Hoses

10 years

NO

Pitot and Static System Hoses

10 years

NO

Vacuum Relief/Regulator Valve Filter
(If Installed)

500 hours

NO

Engine Compartment Flexible FluidCarrying Teflon Hoses (CessnaInstalled) Except Drain Hoses
(Drain hoses are replaced
on condition)

10 years or at engine overhaul,
whichever occurs first
(Note 1)

NO

Engine Compartment Flexible FluidCarrying Rubber Hoses (CessnaInstalled) Except Drain Hoses
(Drain hoses are replaced
on condition)

5 years or at engine overhaul,
whichever occurs first
(Note 1)

NO

Engine Air Filter

500 hours or 36 months,
whichever occurs first
(Note 9)

NO

Engine Mixture, Throttle, and
Propeller Controls

At engine TBO

NO

Check Valve (Turbocharger
Oil Line Check Valve)

Every 1,000 hours of
operation
(Note 10)

NO

Oxygen Bottle - Lightweight Steel
(ICC-3HT, DOT-3HT)

Every 24 years or 4380 cycles,
whichever occurs first

NO

Oxygen Bottle - Composite
(DOT-E8162)

Every 15 years

NO

Engine-Driven Dry Vacuum Pump
Drive Coupling
(Not lubricated with engine oil)

6 years or at vacuum
pump replacement,
whichever occurs first

NO

Engine-Driven Dry Vacuum Pump
(Not lubricated with engine oil)

500 hours
(Note 11)

NO

Standby Dry Vacuum Pump

500 hours or 10 years,
whichever occurs first
(Note 11)

NO

D2007-3-13 Temporary Revision Number 5 - Jan 6/2003
0 Cessna Aircraft Company

3.

Supplier-Established Replacement Time Limits
A.

The following component time limits have been established by specific suppliers and are
reproduced as follows:
Table 2: Supplier-Established Replacement Time Limits
COMPONENT

REPLACEMENT
TIME

OVERHAUL

ELT Battery

(Note 3)

NO

Vacuum Manifold

(Note 4)

NO

Magnetos

(Note 5)

YES

Engine

(Note 6)

YES

Engine Flexible Hoses
(TCM Installed)

(Note 2)

NO

Auxiliary Electric Fuel Pump

(Note 7)

YES

Propeller

(Note 8)

YES

NOTES:
Note 1: This life limit is not intended to allow flexible fluid-carrying Teflon or rubber hoses in a deteriorated or
damaged condition to remain in service.
Replace engine compartment flexible Teflon
(AE3663819BXXXX series hose) fluid-carrying hoses (Cessna-installed only) every ten years or at
engine overhaul, whichever occurs first. Replace engine compartment flexible rubber fluid-carrying
hoses (Cessna-installed only) every five years or at engine overhaul, whichever occurs first (this does
not include drain hoses). Hoses which are beyond these limits and are in a serviceable condition,
must be placed on order immediately and then be replaced within 120 days after receiving the new
hose from Cessna.
Note 2: Refer to Teledyne Continental Service Bulletin SB97-6, or latest revision.
Note 3: Refer to FAR 91.207 for battery replacement time limits.
Note 4: Refer to Airborne Air & Fuel Product Reference Memo No. 39, or latest revision, for replacement
time limits.
Note 5: For airplanes equipped with Slick magnetos, refer to Slick Service Bulletin SB2-80C, or latest
revision, for time limits.
For airplanes equipped with TCM/Bendix magnetos refer to Teledyne Continental Motors Service
Bulletin No. 643, or latest revision, for time limits.
Note 6: Refer to Teledyne Continental Service Information Letter SIL98-9, or latest revision, for time limits.
Note 7: Refer to Cessna Service Bulletin SEB94-7 Revision 1/Dukes Inc. Service Bulletin NO. 0003, or
latest revision.
Note 8: Refer to the applicable McCauley Service Bulletins and Overhaul Manual for replacement and
overhaul information.

Temporary Revision Number 5
6 January 2003

© 2003 CESSNA AIRCRAFT COMPANY

2-29

1

Note 9: The air filter may be cleaned, refer to Section 2 of this service manual and for airplanes equipped
with an air filter manufactured by Donaldson, Refer to Donaldson Aircraft Filters Service
Instructions P46-9075 for detailed servicing instructions.
The address for Donaldson Aircraft Filters is:
Customer Service
115 E. Steels Comers RD
Stow OH. 44224
Do not overservice the air filter, overservicing increases the risk of damage to the air filter from
excessive handling. A damaged/worn air filter may expose the engine to unfiltered air and result in
damage/excessive wear to the engine.
Note 10: Replace the turbocharger oil line check valve every 1,000 hours of operation (Refer to Cessna Service
Bulletin SEB91-7 Revision 1, or latest revision).
Note 11: Replace engine driven dry vacuum pump not equipped with a wear indicator every 500 hours of
operation, or replace according to the vacuum pump manufacturer's recommended inspection and
replacement interval, whichever occurs first.
Replace standby vacuum pump not equipped with a wear indicator every 500 hours of operation or
10 years, whichever occurs first, or replace according to the vacuum pump manufacturer's
recommended inspection and replacement interval, whichever occurs first.
For a vacuum pump equipped with a wear indicator, replace pump according to the vacuum pump
manufacturer's recommended inspection and replacement intervals.

1 2-30

© 2003 CESSNA AIRCRAFT COMPANY

Temporary Revision Number 5
6 January 2003

SECTION 3
FUSE LAGE
TABLE OF CONTENTS

Page

. 3-1
. ....
FUSELAGE ........
3-1
Windshield and Windows .........
3-1
Description .............
3-1
............
Cleaning
3-1
Waxing ................
3-1
Repairs ..............
3-1
.............
Scratches
..
. 3-2
..........
Cracks
3-4
Windshield ..............
3-4
Removal and Installation .......
3-4
W indows . . . . . .. . . . . . . . .
3-4
Movable, Fixed and Rear.......
3-4
Cabin Doors .............
3-4
Removal and Installation .......
3-4
Adjustment .............
3-4
Weatherstrip ............
..
3-4
Wedge Adjustment ...
3-4
Cabin Door Latches .........
3-4
Description ...........
3-4
Adjustment
...........
...
3-4
............
Lock.
3-5
Indexing Inside Handle ........
3-5
Assist Straps .............
Removal and Installation .......
3-5
3-5
..............
Baggage Door
3-5
Removal and Installation .......
3-5
Cargo Doors ..............
3-5
............
Description
3-5
Removal and Installation .......
3-5
..
Latches ....
Removal and Installation .
.
. 3-5
3-5
Rigging
.............
3-10
Seats . . . . . . . ..
3-10
Pilot and Copilot ...........
. 3-10
..
Reclining Back ........
3-10
Reclining Back/Vertical Adjust ....
3-1.

FUSELAGE.

3-2.

WINDSHIELD AND WINDOWS.

3-3. DESCRIPTION. The windshield and windows
are single-piece acrylic plastic panels set in sealing
strips and held by formed retaining strips secured
to the fuselage with screws and rivets. Presstite No.
579.6 sealing compound used in conjunction with a
felt seal is applied to all edges of the windshield and
windows with the exception of the wing root area.
The wing root fairing has a heavy felt strip that completes the windshield sealing.
3-4.

CLEANING.

(Refer to Section 2.)

3-5. WAXING. Waxing will fill in minor scratches
in clear plastic and help protect the surface from
further abrasion. Use a good grade of commercial
wax applied in a thin, even coat. Bring the wax to a
high polish by rubbing lightly with a clean, dry flannel cloth.
3-6.

REPAIRS.

Damaged window panels and wind-

Articulating Recline/Vertical
Adjust ..............
Description . ...........
Removal and Installation .......
...........
Center and Rear
Reclining Back/Fore-and-Aft
....
Adjust ........
Non-Reclining Back/Fore-and-Aft
Adjust ..............
Description ............
Repair . . . . . . . . . . . . . .
. .
Cabin Upholstery ........
Materials and Tools .........
........
Soundproofing ....
Cabin Headliner ...........
............
Removal ..
. ........
Installation .....
.
Upholstery Side Panels ......
Windlace (Door Seal). ..........
Carpeting
..............
Safety Provisions ............
Cargo Tie-Downs ...........
Safety Belts ............
Shoulder Harness ...........
. ..
Glider Tow Hook ......
.
Rear View Mirror ........
..............
Cargo Pack
....
Removal ........
Installation ............
Cowl Flap Baffles and Control Extensions
Removal ............
...........
Installation .
Casket Carrier . . . . . . . . . . . .
Description .............
Installation .............
... ..
.
Removal . ....

3-10
3-10
3-10
3-10
3-10

.
.
.

.
.
.
.
.
.
.
.

3-10
3-10
3-10
3-10
3-10
3-10
3-10
3-10
3-10
3-23
3-23
3-23
3-23
3-23
3-23
3-24
3-24
3-24
3-24
3-24
3-24
3-24
3-25
3-25
3-25
3-25
3-25
3-29

shield may be removed and replaced if damage is
extensive. However, certain repairs as prescribed
in the following paragraphs can be made successfully
without removing damaged part from aircraft. Three
types of temporary repairs for cracked plastic are
possible. No repairs of any kind are recommended
on highly-stressed or compound curves where repair
would be likely to affect pilot's field of vision.
Curved areas are more difficult to repair than flat
areas and any replaced area is both structurally and
optically inferior to the original surface.
3-7. SCRATCHES. Scratches on clear plastic surfaces can be removed by hand-sanding operations
followed by buffing and polishing, if steps below are
followed carefully.
a. Wrap a piece of No. 320 (or finer) sandpaper or
abrasive cloth around a rubber pad or wood block.
Rub surface around scratch with a circular motion,
keeping abrasive constantly wet with clean water to
prevent scratching surface further. Use minimum
pressure and cover an area large enough to prevent
formation of "bull's-eyes" or other optical distortions.
Change 2

3-1

Figure 3-1.

Repair of Windshield and Windows

CAUTION
Do not use a coarse grade of abrasive.
320 is of maximum coarseness.

No.

b. Continue sanding operation, using progressively
finer grade abrasives until scratches disappear.
c. When scratches have been removed, wash area
thoroughly with clean water to remove all gritty partides. The entire sanded area will be clouded with
minute scratches which must be removed to restore
transparency.
d. Apply fresh tallow or buffing compound to a
motor-driven buffing wheel. Hold wheel against plastic surface, moving it constantly over damaged area
until cloudy appearance disappears. A 2000-foot-perminute surface speed is recommended to prevent
overheating and distortion. (Example: 750 rpm
polishing machine with a 10 inch buffing bonnet.)
NOTE
Polishing can be accomplished by hand but
will require a considerably longer period
of time to attain the same result as produced by a buffing wheel.
e. When buffing is finished, wash area thoroughly
and dry with a soft flannel cloth. Allow surface to
cool and inspect area to determine if full transparency has been restored. Apply a thin coat of hard
wax and polish surface lightly with a clean flannel
cloth.
NOTE
Rubbing plastic surface with a dry cloth

3-2

will build up an electrostatic charge which
attracts dirt particles and may eventually
cause
scratching of surface. After wax
has hardened, dissipate this charge by rubbing surface with a slightly damp chamois.
This will also remove dust particles which
have collected while wax is hardening.
f. Minute hairline scratches can often be removed
by rubbing with commercial automobile body cleaner or fine-grade rubbing compound. Apply with a
soft, clean, dry cloth or imitation chamois.
3-8. CRACKS. (Refer to figure 3-1.)
a. When a crack appears, drill a hole at end of
crack to prevent further spreading. Hole should be
approximately 1/8 inch in diameter, depending on
length of crack and thickness of material.
b. Temporary repairs to flat surfaces can be accomplished by placing a thin strip of wood over each
side of surface and inserting small bolts through
wood and plastic. A cushion of sheet rubber or aircraft fabric should be placed between wood and plastic on both sides.
c. A temporary repair can be made on a curved surface by placing fabric patches over affected areas.
Secure patches with aircraft dope. Specification No.
MIL-D-5549, or lacquer. Specification No. MIL-L7178. Lacquer thinner, Specification No. MIL-T6094 can also be used to secure patch.
d. A temporary repair can be made by drilling
small holes along both sides of crack 1/4 to 1/8 inch
apart and lacing edges together with soft wire.
Small-stranded antenna wire makes a good temporary
lacing material. This type of repair is used as a
temporary measure ONLY, and as soon as facilities
are available, panel should be replaced.

B

C

A

Detail

A

,k
^^

BEGINNING WITH
SERIAL U20603021

2

Detail C

Felt Seal
Retainer
3. Window
4 Fuselage Skin
5. Window Frame
6. Window
7. Latch Assembly
8. Stop
1.
2.

9.
10.
11.
12.

Detail

Fuselage Structure
Hinge
Striker Plate
Spring

Figure 3-2.

D
'
4

(

Typical Side Window Seals
NOTE
Presstite No. 579.6 sealer should be applied
to all edges of windshield and windows where
felt sealing strip (1) is used.

Windshield and Window Installation.
Change 3

3-3

3-9.

WINDSHIELD.

(Refer to figure 3-2.)

3-10. REMOVAL AND INSTALLATION.
a. Drill out rivets securing top retainer strip.
b. Remove screws securing front retainer strip.
c. Remove wing fairings over windshield edges.
d. Pull windshield straight forward, out of side retainers.
e. Reverse preceding steps for reinstallation. Apply
felt strip and sealing compound to all edges of windshield to prevent leaks. Check fit and carefully file
or grind away excess plastic.
3-11.

WINDOWS.

3-12.
MOVABLE. (Refer to figure 3-2 ) A movable
window hinged at the top is installed in the left cabin
door thru 1975 models and beginning with 1976 models
in the RH forward side window position. The window
assembly is a tinted plastic and frame unit which may
be replaced by removing hinge pins and disconnecting
window stop. To remove plastic panel from frame,
drill out blind rivets at frame splice. When replacing
plastic panel, ensure an adequate coating of Presstite
579.6 sealing compound is applied to all edges of panel.
3-13. FIXED. (Refer to figure 3-2.) Fixed windows, mounted in sealing strips and sealing compound, are held in place by various retainer strips.
To replace side windows, remove upholstery and
trim panels as necessary and drill out rivets securing retainers.
3-14. REAR. (Refer to figure 3-2.) The curved triangular rear side windows are mounted in retaining
and sealing strips. Windows are removed from inside the cabin after rivets securing strips are drilled
out. Removal of the rectangular rear window requires
drilling out three rows of rivets immediately forward
and above the window. Remove screws securing retainer strips at each side of the window and deflect
strips up and aft from skin splice above the window.
Remove the window from inside the aircraft. Reverse
the preceding procedure for installation. Check fit
of the new window and carefully file or grind away
excess plastic. Apply felt strips and sealing compond
to all edges.
3-15.

CABIN DOORS.

(Refer to figure 3-3.)

3-16 REMOVAL AND INSTALLATION. Removal
of cabin doors is accomplished by removing screws
which attach hinges and door stop or by removing
hinge pins attaching door and door stop. If permanent hinge pins are removed from door hinges, they
may be replaced by clevis pins secured with cotter
pins or new hinge pins may be installed and "spinbradded." When fitting a new door, some trimming
of door skin at edges and some forming of door edges
with a soft mallet may be necessary to achieve a
good fit. Forming of the flanges on the bonded door
is not permissible as forming of the flanges could
cause damage to the bonded area.
3-17. ADJUSTMENT. Cabin doors should be adjusted so skin fairs with fuselage skin. Slots at
latch plate permit repositioning of striker plate.
3-4

Change 3

Depth of latch engagement may be changed by adding
or removing washers or shims between striker plate
and doorpost.
3-18. WEATHERSTRIP.
Rubber seals are installed
around the edges of the cabin door. Beginning with
serial U20602790 an improved type door seal is used
which has a hollow center and small flutes extending
along its length. When replacing door seals ensure
mating surfaces are clean, dry and free of oil and
grease. Position butt ends of seal at door low point
and cut a small notch in the hollow seal for drainage.
Apply a thin, even coat of EC-880 adhesive ( 3M Co)
or equivalent to each surface and allow to dry until
tacky before pressing into place.
3-19. WEDGE ADJUSTMENT. Wedges at upper
forward edge of door aid in preventing air leaks
at this point. They engage as door is closed. Several attaching holes are located in wedges and holes
which gives best results should be selected.
3-20.
3-6.)

CABIN DOOR LATCHES.

(Refer to figure

3-21. DESCRIPTION. The cabin door latch is a
push-pull bolt type, utilizing a rotary clutch for positive bolt engagement. As door is closed, teeth on
underside of bolt engage gear teeth on clutch. The
clutch gear rotates in one direction only and holds
door until handle is moved to LOCK position, driving
bolt into slot.
3-22. ADJUSTMENT. Adjustment of latch or clutch
cover is afforded by oversize and/or slotted holes.
This adjustment ensures sufficient gear-to-bolt engagement and proper alignment. To adjust bolt (item 2)
figure 3-6. loosen the four latch base bolts (item 29)
sufficient to move latch base plate aft to extend the bolt
or forward to retract the bolt.
{CAUTION
Close the door carefully alter adjustment
and check for clearance between door jamb
and bolt and alignment with clutch assembly.
NOTE
Lubricate door latch per Section 2. No lubrication is recommended for rotary clutch.
3-23. LOCK. In addition to interior locks, a cylinder and key type lock is installed on left door. If
lock is to be replaced, the new one may be modified
to accept original key. This is desirable, as the
same key is used for ignition switch and cabin door
lock. After removing old lock from door, proceed
as follows:
a. Remove lock cylinder from new housing.
b. Insert original key into new cylinder and file off
any protruding tumblers flush with cylinder. Without
removing key, check that cylinder rotates freely in
housing.
c. Install lock assembly in door and check lock
operation with door open.
d. Destroy new key and disregard code number on
cylinder.

3-24. INDEXING INSIDE HANDLE. (Refer to figure
3-6.) When inside door handleis removed, install
in relation to position of bolt (2) which is spring-loaded to CLOSE position. The following procedure may
be used:
a. THRU SERIALS P20600647 AND U20602199.
(Refer to figure 3-6, sheet 1.)
1. Temporarily install handle (15) on shaft
assembly (19) approximately vertical.
2. Move handle (15) back and forth until handle
centers in spring-loaded position.
3. Without rotating shaft assembly (19), remove
handle and install spring (9) and escutcheon (13).
4. Install handle (15) in vertical position and
install clip (16).
5. Ensure bolt (2) clears doorpost and teeth engage clutch gear (26) when handle (15) is in CLOSE
position.
b. BEGINNING WITH SERIALS U20602200. (Refer to
figure 3-6, sheet 2.) These models feature an inside
door handle positioned forward on the door. The handle
folds into the armrest when in the "LOCKED" position.
1. Complete steps 1 and 2 as outlined in step

used to hold doors open. An entrance step is located
on fuselage, below front cargo door. Flight with
doors removed is only permissible when an optional
spoiler kit is installed. This spoiler kit consists of
a spoiler assembly which attaches to front door hinge
points and deflects air away from door opening.
Addition of screws to rear wall is required with installation of spoiler kit.
NOTE
A flap interrupt switch is installed to prevent
operation of flaps with cargo doors open.
Switch adjustment is provided by means of
slotted holes on front cargo door frame. A
switch depressor is provided with spoiler kit
to retain use of flaps.
3-29. REMOVAL AND INSTALLATION.
a. Remove cotter pins and hinge pins from door
hinges.
b. Disconnect door stops from doors.
c. Reverse preceding steps for installation.

,a. "
2. Without rotating shaft assembly (19), remove
handle and install spring (9) and nylon washer (10).
3. Install handle (15) to align with CLOSE position on upholstery panel (12) .
4. Complete step "5" as outlined in step "a."
5. Readjust handle on serrated shaft as necessary to position the forward end of the handle approx.
8° above the handle shaft centerline when in the LOCKED position.
1.
3-24A. ASSIST STRAPS. (Refer to figure 3-3A)
3-24B. REMOVAL AND INSTALLATION. Figure
3-3A may be used as a guide for removal and
installation of the assist straps.
3-25.

BAGGAGE DOOR.

CARGO DOORS.

(Refer to figure 3-5.)

3-28. DESCRIPTION. U206 and TU206 aircraft are
equipped with two cargo doors located on the right
side of fuselage. The aft door is hinged at fuselage
station 112 and is a structural, load-carrying member when closed and locked. The aft door handle is
located in forward edge of door and is inaccessible
with forward door closed, preventing inadvertent
opening during flight. As rear door handle is moved
to CLOSED position, hooks engage latch plates on
upper and lower door sills holding door tightly closed.
Telescoping door stops, with detent positions, are

LATCHES.

(Refer to figures 3-5 and 3-6.)

3-31. REMOVAL AND INSTALLATION. Figures
3-5 and 3-6 show details of cargo door latches and
may be used as guides during removal, disassembly,
assembly and installation.
3-32. RIGGING. (Refer to figure 3-5.)
Three results must be obtained
a. Three results must be obtained by rigging.
Hooks (8) must fully engage latch plates (3),
but must clear them .05" minimum as door is opened.
2. Load-carrying pins (7) must fully engage
their sockets when door is locked.
3. Door must be flush with fuselage skin when
door is locked.
NOTE

(Refer to figure 3-4.)

3-26. REMOVAL AND INSTALLATION.
a. Disconnect door stop (2) at door.
b. Remove hinge pins (3) securing door to hinges
(4).
c. Reverse preceding steps for installation.
3-27.

3-30.

Adjusting door slightly less than flush is permissible if air leaks around door seal are encountered.
There are four sets of adjustments for rigging:
1. Adjusting bolts (10). These determine depth
of hook engagement and clearance of hooks as door is
opened.
2. Slots in latch plates (3). Plates may be
moved inboard or outboard as necessary for full
load-carrying pin engagement.
3. Washers under socket (6). These may be
added as required to make door flush with fuselage
skins.
4. Turnbuckles (11). These must be adjusted
to cause both hooks to pull door closed tightly. Handle should snap over-center snugly, but excessive
force should not be required for handle operation.
b.

Change 3

3-5

NOTE

NOTE

Spray cabin and window seals
with MS-122 (Miller-Stephenson
Chem. Corp., Danbury, Conn.)
or equivalent. Caution, do not
overspray; confine to seals.

Forming of the flanges on the
bonded door is not permissible
as forming could cause damage
to the bonded area.

7

SEE FIGURE 3-6

Detail A

Detail

B

A

23
DOOR INSTALLATION
THRU AIRCRAFT

DetailC

Detail D

19

THRU 1972

4

15-'\ \ ^?

!

'20
22 j3.
^ ^
'

N~22

Da

DD

"p

i5.
:> ^8.

2-^9-F9.

Detail
2
BEGINNINGWITH
AIRCRAFT SERIAL
AIRCRAFT
SERIAL
TU20601875

Change 3

1. Upholstery Clip
2. Upholstery Panel
Wedge
4. Spring
Window Stop
6. Window Hinge
7. Latch Plate
Cabin Door
Window Frame
10. Window
11. Washer
12. Nut

.25.
Figure 3-3.

3-6

14

5

D

24

Cabin Door Installation (Sheet 1 of 2).

13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.

Lock Assembly
Latch Assembly
Door Stop Arm
Spring-Loaded Plunger
Wedge
Spacer
Stop Assembly
Reinforcement
Hinge
Pin
Lower Hinge
Upper Hinge
Door Jamb

Refer to Fire

H^'

Detail

/

[

'

,

'1y 24
^^ / 5 /,/^^ ^^

\\et

^is

;/

l,

14

Rotated 180

C

to Figure 3-6.
'
FigureRefer

to^

."i<1;
-^.

DetailE

'-

12C^\

Forming the flange
a bonded door
is not permissable

" Jof

/

AiE

..

as it could cause
material separation.

F

DOOR INSTALLATION
BEGINNING WITH
AIRCRAFT SERIAL
U20602200

1

20

12

t11l

2

Detail G

NOTE
Spray cabin and window seals
with MS-122 (Miller-Stephenson
Chem. Corp., Danbury, Conn.)
or equivalent. Caution, do not
overspray; confine to seals

Detail

F

Rotated 180

BEGINNING WITH
AIRCRAFT SERIAL
SERIAL
ARRF
75
U0

Figure 3-3. Cabin Door Installation (Sheet 2 of 2)
Change 3

3-6A

Detail A
DetailC

AVAILABLE BEGINNING WITH
AIRCRAFT SERIAL U20602580

BEGINNING WITH AIRCRAFT
SERIAL U20602200

1.

Screw

2. Pull Handle

3.
4.
5.
6.

Clamp Cover
Clamp
Fuselage
Window Moulding

\
,
/

h.,
,

,

7. Door Post

Detail B
BEGINNING WITH AIRCRAFT
SERIAL U20602360
3-6B

Change 2

3-6B

Change 2

Figure 3-3A. Assist Strap Installation

2

A

ROTATED 180°

1. Window

2. Stop Arm Attach

6

10
Detail

A

9.
*Use as required to align
handle flush with outside
skin.

Figure 3-4.

3. Hinge Pin
4. Hinge
5. Upholstery Panel
6. Cam
7. Latch Assembly
8. Lock
Shim
10. Handle
11. Baggage Door Skin

Baggage Door Installation
Change 3

3-7

2

FWD 3

SEE FIGURE 3-6

11

Detail A

11

/

,

-

$1

Ir^

~~~

~

e'

lDetail

B

NOTE
Sockets (6) are mounted in the upper a
and lower door sills. Install an
abrasive shim beneath latch plate (3)
to prevent latch plate from slipping.

Detail C

1. Door Stop
2. Flap Interrupt Switch
3. Latch Plate
4. Nut
5. Washer
6. Socket
7. Load Carrying Pin

8.
9.
10.
11.
12.
13.
14.

Upper Hook
Upper Latch Carrier
Adjusting Bolt
Turnbuckle
Bushing
Handle
Cover
Figure 3-5.

3-8

Change 3

To aid in cargo loading, the center
seat bolt attach points on the floor
are designed to fold flat.
A tee handle is stowed in the glove
compartment. The front cargo
door may be locked and unlocked
externally through a hole opposite
the inside handle.

Cargo Door Installation

**THRU AIRCRAFT SERIALS U206-1444
AND P20600603
*THRU AIRCRAFT SERIALS U20601874
AND P20600648
*AIRCRAFT SERIALS P20600604 THRU
P20600648 AND U20601445 THRU
U20601587
*BEGINNING WITH AIRCRAFT SERIALS
U20601445 AND P20600604
**BEGINNING WITH AIRCRAFT SERIAL
U20601588
**BEGINNING
U20601875

WITH AIRCRAFT SERIAL

ROTATED 90°

Top

Rotary clutch components
are matched upon assembly.
The clutch mechanism, if
defective, should be replaced
as a unit.

Bolt
1. Guide
2. Bolt
3. Side Bolt Guide
4. Base Bolt Guide
5. Latch Base Plate
6. Abrasive Pad
7. Lockplate
8. Bracket
9. Spring
Nylon Washer
10.
11. Placard
12. Upholstery Panel

13. Escutcheon
14. Placard
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.

CARGO DOOR
ROTARY CLUTCH

25.
26.
27.
28.
29.
30.
Figure 3-6.

Inside Handle
Clip
Plate Assembly
Support
Shaft Assembly
Bolt Push Rod
Outside Handle
Pull Bar
Mounting Structure
Shim
Rotary Clutch
Guide
Door Post
Cover
Adjustment Screw
Push-Pull Rod

Door Latch and Rotary Clutch Components (Sheet 1 of 2)
Change 2

3-8A/3-8B(blank)

NOTE

Refer
to paragraph
3-22 for
bolt (item 2)

A

20
Detail A

21
20

23

ROTARY CLUTCH

SERIAL
Set adjustment screw (29)
in the slot to maintain door
handle 8- 15' above center
line of handle shaft when the
door is in the locked position.
Figure 3-6.

20602810
17

BEGINNING WITH AIRCRAFT
SERIAL U20602200

Rotary clutch components are
matched upon assembly. The
clutch mechanism, if defective, should be replaced as a
unit.

Door Latch and Rotary Clutch Components (Sheet 2 of 2)
Change 3

3-9

3-33.

SEATS.

(Refer to figure 3-7.)

3-34.

PILOT AND COPILOT.
a. RECLINING BACK. (Standard pilot/Optional copilot.)
b. RECLINING BACK/VERTICAL ADJUST.
(Optional 1969 ONLY. )
c. ARTICULATING RECLINE/VERTICAL
ADJUST. (Optional 1970 AND ON.)
3-35. DESCRIPTION. These seats are manuallyoperated throughout their full range of operation.
Seat stops are provided to limit fore-and-aft travel.
Install seat stops on rails as follows:
1. Pilots seat: inbd rail fwd and aft.
2. Copilots seat: outbd rail fwd and aft.
3. Center L H seat: outbd rail fwd and aft.
4. Center R H seat: outbd rail fwd and inbd rail aft.
5 Aft L H seat: outbd rail fwd and aft.
6. Aft R H seat: outbd rail aft only.
3-36. REMOVAL AND INSTALLATION.
a. Remove seat stops from rails.
b. Slide seat fore-and-aft to disengage seat rollers
from rails.
c. Lift seat out.
d. Reverse the preceding steps for installation.
Ensure all seat stops are reinstalled.
WAR NI
NG
It is extremely important that pilot's seat
stops are installed, since acceleration and
deceleration could possiblypermit seat
to become disengaged from seat rails and
create a hazardous situation, especially during take-off and landing.
3-37.

by a mechanic unfamiliar with upholstery practices.
the mechanic should make careful notes during removal of each item to facilitate its replacement later.
3-41. MATERIALS AND TOOLS. Materials and
tools will vary with job. Scissors for trimming upholstery to size and a dull-bladed putty knife for
wedging material beneath retainer strips are the
only tools required for most trim work. Use industrial rubber cement to hold soundproofing mats
and fabric edges in place. Refer to Section 18 for
thermo-plastic repairs.
3-42. SOUNDPROOFING. The aircraft is insulated
with spun glass mat-type insulation and a sound deadener compound applied to inner surfaces of skin in
most areas of cabin and baggage compartment.
All soundproofing material should be replaced
in its original position any time it is removed.
A soundproofing panel is placed in the gap between the wing and fuselage and held in place
by the wing root fairing.
3-43.

CABIN HEADLINER.

(Refer to figure 3-10.)

3-44. REMOVAL.
a. Remove sun visors, all inside finish strips and
plates, door post upper shields, front spar trim
shield, dome light console and any other visible retainers securing headliner.
b. Work edges of headliner free from metal teeth
which hold fabric.
c. Starting at front of headliner, work headliner
down, removing screws through metal tabs which
hold wire bows to cabin top. Pry loose outer ends
of bows from retainers above doors. Detach each
bow in succession.

CENTER AND REAR.
a. RECLINING BACK/FORE-AND-AFT AD-

NOTE

.JUST.
b.
ADJUST.

NON-RECLINING BACK/FORE-AND-AFT

3-38. DESCRIPTION. These seats are provided
with fore-and-aft adjustment provisions. Seat stops
are installed to limit travel. Removal and installation is outlined in paragraph 3-36.
3-39. REPAIR. Replacement of defective parts is
recommended in repair of seats. However, a
cracked framework may be welded, provided crack
is not in an area of stress concentration (close to a
hinge or bearing point). The square-tube framework
is 6061 aluminum, heat-treated to a T-6 condition.
Use a heliarc weld on these seats, as torch welds
will destroy heat-treatment of frame structure. Figure 3-8 outlines instructions for replacing defective
cams on reclining seat backs.
3-40. CABIN UPHOLSTERY. Due to the wide selection of fabrics, styles and colors, it is impossible to
depict each particular type of upholstery. The following paragraphs describe general procedures which
will serve as a guide in removal and replacement of
upholstery. Major work, if possible, should be done
by an experienced mechanic. If work must be done
3-10

Change 1

Always work from front to rear when removing headliner.
d. Remove headliner assembly and bows from aircraft.
NOTE
Due to difference in length and contour of
wire bows, each bow should be tagged to
assure proper location in headliner.
e.

Remove spun glass soundproofing panels.
NOTE
The lightweight soundproofing panels are
held in place with industrial rubber cement.

3-45. INSTALLATION.
a. Before installation, check all items concealed
by headliner for security. Use wide cloth tape to
secure loose wires to fuselage and to seal openings
in wing roots. Straighten tabs bent during removal
of headliner.

PILOT AND COPILOT SEATS

THRU 1972

1.

Recline Handle

12

2. Pin
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.

RECLINING BACK/
VERTICAL ADJUST
VERTICAL
(OPTIONAL 1969 ONLY)

Shaft
Seat Bottom
Seat Back
Bushing
Spacer
Spring
Seat Adjustment Pawl
Seat Roller
Bracket
Washer
Adjustment Pin
Fore/Aft Adjustment Handle
Seat Stop
Channel
Bellcrank
Vertical Adjustment Handle
Adjustment Screw
Seat Structure
Torque Tube

NOTE
Seat back cams are similar for both
seats illustrated. Refer to figure 3-8
for replacement.

Figure 3-7.

14

Seat Installation(Sheet I of 11)

|

Change 1

3-11

PILOT AND COPILOT SEAT
(STANDARD BEGINNING WITH 1973)

3

1

2

RECLINING BACK

1

BEGINNING WITH SERIAL U20603021
9

12

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

Recline Handle
Pin
Link Assembly
Torque Tube
Seat Back
Recline Cam
Bushing
Spacer
Spring
Pawl
Roller
Adjustment Pin
Fore/Aft Adjustment Handle

14. Seat Bottom
15. Seat Belt Retainer
Figure 3-7.
3-12

Change 3

Seat Installation (Sheet 2 of 11)

PILOT AND COPILOT SEAT

ARTICULATING RECLINE/
VERTICAL ADJUST
OPTIONAL 1970 THRU
1972)

Detail B
1.
2.
3.
4.
5.

14

Vertical Adjustment Handle
Adjustment Pin
Fore-and-Aft Adjustment Handle
Seat Bottom
Articulating Adjustment Handle
Figure 3-7.

6. Bellcrank
7. Adjustment Screw
8. Seat Back
9. Magazine Pocket

10.
11.
12.
13.
14.

Trim Bracket
Channel
Torque Tube
Seat Structure
Roller

Seat Installation(Sheet 3 of 11)
Change 1

3-13

PILOT AND COPILOT SEAT
(OPTIONAL THRU 1973)

VERTICAL ADJUST

Detail

4

A

8

,>Mi

<

,,

Detail B

14
1. Vertical Adjustment Handle
2.
3.
4.
5.
6.

8. Bellcrank

Fore/Aft Adjustment Handle
Adjustment Pin
Spring
Seat Bottom
Articulating Adjustment Handle

7. Adjustment Screw
*

9.
10.
11.
12.
13.

14. Roller
Figure 3-7.

3-14

Change 1

Seat Back
Spacer
Channel
Torque Tube
Seat Structure

Seat Installation (Sheet 4 of 11)

PILOT AND COPILOT SEAT
BEGINNING WITH 1974 MODELS
(OPTIONAL INSTALLATION)

9

9

ARTICULATING BACK/
VERTICAL ADJUST

Detail A

1. Vertical Adjustment Handle

9. Seat
Sack
SERIALS
U20602410
THRU U20603020
* BEGINNING WITH SERIALS U20603021

Detail

1. Vertical Adjustment Handle
9. Seat Back
2. Fore/Art Adjustment Handle
10.

Spacer

3.
4.
5.
6.
7.
8.

Channel
Torque Tube
Seat Structure
Roller
Stiffner
Seat Belt Retainer

Adjustment Pin
Spring
Seat Bottom
Articulating Adjustment Handle
Adjustment Screw
Bellcrank

11.
12.
13.
14.
15.
16.

Figure 3-7.

/

B

/14

Seat Installation (Sheet 5 of 11).
Change 3

3-14A/3-14B(blank)

CENTER
1. Reclining Adjustment Handle
2. Spacer
3. Seat Bottom
4. Torque Tube
5. Link
6. Bellcrank
7. Fore/Aft Adjustment Handle
8. Fore/Aft Adjustment Pin
9. Spring
10. Spring Positioning Support
11. Reclining Adjustment Pawl
12. Bushing
13. Bushing
14. Seat Back
1

206 SERIES SERIALS

2
P206-0520AND ON
20601588 AND ON

*RIGHT HAND SEAT ONLY

AIRCRAFT SERIALS
U2060158R AND ON

0,

/\S

^

\\1\

^134

2*

.

SEAA
LEFT HAND
SEAT ONLY\

\
RIGHT HAND
SEAT ONLY
Detail

9
19*

A

'<^-^~~

\

Detail

B

9

Figure 3-7.

Seat Installation (Sheet 6 of 11)
Change 1

3-15

1.
2.
3.
4.

Seat Stop
Fore/Aft Adjustment Handle
Adjustment Pin
Roller

Figure 3-7. Seat Installation (Sheet 7 of 11)
3-16

Change 1

REAR
1969
P206 AND TP206

4

e

DetailA

il B

Detail

Detail

11

1.
2.
3.
4.

5. Seat Back
6. Bushing
7. Washer
8. Pin

Roller
Seat Bottom
Pawl
Spring

Figure 3-7.

9.
10.
11.
12.

C

3

Handle
Spring
Adjustment Pin
Handle

I
Seat Installation (Sheet 8 of 11)
Change 1

3-17

REAR
1970
P206 AND TP206

4

3

Detail A

2. Seat Bottom
3. Pawl
4. Spring
5. Seat Back
6. Bushing
7. Washer
8. Pin
9. Reclining Adjustment Handle
10. Adjustment Arm
11. Spacer

12.

Fore/Aft Adjustment Handle

13.

Adjustment Pin
Figure 3-7.

3-18

Change 1

Detail

D

Seat Installation (Sheet 9 of 11)

1.
2.
3.
4.
5.
6.

Roller
Seat Bottom
Seat Back
Spring
Adjustment Pin
Fore/Aft Adjustment Handle

Figure 3-7.

Seat Installation (Sheet 10 of 11)

|

Change 1

3-19

REAR
1970
U206 AND TU206

B

3.
4.

Seat Back
Bushing

7.

Figure 3-7.
3-20

Change 1

11

Spacer

Seat Installation (Sheet 11 of 11)

Detail A

10.
11.

Spring
Adjustment Pin

CLEVIS BOLT (REF)

SEAT BACK (REF)

2.50" R. (CONSTANT AT EACH NOTCH)

REPLACEMENT CAM:
1414230-1 (SINGLE
ADJUSTABLE SEAT)

PAWL (REF)

ADJUSTABLE
1414111-5 (VERTICALLY
SEAT

ADJUSTABLE SEAT)

REPLACEMENT PROCEDURE:
a.

Remove seat from aircraft.

b.

Remove plastic upholstery panels from aft side of seat back, then loosen upholstery retaining
rings and upholstery material as required to expose the rivets retaining the old cam assembly.

c.

Drill out existing rivets and insert new cam assembly (2).
engages first cam slot as shown.

d.

Position the cam so each slot bottom aligns with the 2. 50" radius as shown.

e.

Clamp securely in this position and check travel of cam. Pawl must contact bottom of each cam
slot. Using existing holes in seat frame, drill through new cam and secure with MS20470AD6
rivets.

Position seat back so that pawl (3)

f. Reinstall upholstery, upholstery panels and seat.

Figure 3-8.

Seat Back Cam Replacement
3-21

1.
2.
3.
4.
5.
6.
7.
8.

Hook
Screw
Shoulder Harness
Cover
Clip
Shoulder Harness
Spacer
Washer

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

Bolt
Cup
Seat Belt
Latch Assembly
Eye Bolt
Nut
Bracket
Fitting

Figure 3-9.
3-22

Change 3

17.
18.
19.
20.
21.
22.

Quick-Release Belt
Stowage Tray
Inertia Reel
Attaching Plate
Aircraft Structure
Inertia Reel Cover

Seat Belt and Shoulder Harness Installation (Sheet 1 of 3)

K
Detail K
Detail J
Pilot and Copilot
positions only

77

17

Detail

L

Detail M
Required on Australian
Aircraft only.

AIRCRAFT SERIAL
U20602200 THRU U20602379

Figure 3-9.

Seat Belt and Shoulder Harness Installation (Sheet 2 of 3)
Change 3

3-22A

*

N

4

-

Detail N
BEGINNING WITH

U20602380

...--.
.

..

20

-

22

Detail O
INERTIA REEL INSTALLATION
BEGINNING WITH U20602580

Figure 3-9. Seat Belt and Shoulder Harness Installation(Sheet 3 of 3)

3-22B

Change 2

21

1. Soundproofing
2. Zipper
3. Headliner
4. Tiara Strip
5. Trim Shield

Figure 3-10.
b. Apply cement to inside of skin in areas where
soundproofing panels are not supported by wire bows
and press soundproofing in place.
c. Insert wire bows into headliner seams and secure rearmost edges of headliner after positioning
two bows at rear of headliner. Stretch material
along edges to ensure it is properly centered, but do
not stretch enough to destroy ceiling contours or
distort wire bows. Secure edges of headliner with
metal teeth or rubber cement.
d. Work headliner forward, installing each wire
bow in place with tabs. Wedge ends of wire bows
into retainer strips. Stretch headliner just taut
enough to avoid wrinkles and maintain a smooth contour.
e. When all bows are in place and fabric edges are
secured, trim off excess fabric and reinstall all
items removed.
3-46. UPHOLSTERY SIDE PANELS. Removal of
upholstery side panels is accomplished by removing
seats for access, then removing parts attaching
panels. Remove screws, retaining strips, arm
rests and ash trays as required to free panels. Automotive type spring clips attach most door panels. A
dull putty knife makes an excellent tool for prying
loose clips. When installing upholstery side panels,
do not over-tighten sheet metal screws. Larger
screws may be used in enlarged holes as long as
area behind hole is checked for electrical wiring,
fuel lines and other components which might be damaged by using a longer screw.

Cabin Headliner
3-47. WINDLACE (DOOR SEAL). To furnish an
ornamental edging for door opening and to provide
additional sealing, a windlace is installed between
upholstery panels or trim panels and doorpost structure. The windlace is held in place by sheet metal
screws.
3-48. CARPETING. Cabin area and baggage compartment carpeting is held in place by rubber cement, small sheet metal screws and retaining strips.
When fitting a new carpet, use old one as a pattern
for trimming and marking screw holes.

3-49.

SAFETY PROVISIONS.

3-50. CARGO TIE-DOWNS. Cargo tie-downs are
used to ensure baggage cannot enter seating area
during flight. Methods of attaching tie-downs are illustrated in figure 3-11. The eyebolt and nutplate
can be located at various points. The sliding tiedown lug also utilizes eyebolt and attaches to a seat
rail. Different combinations of all four may be used.
3-51. SAFETY BELTS. Safety belts should be replaced if frayed or cut, latches are defective or
stitching is broken. Attaching parts should be replaced if excessively worn or defective. (Refer to
figure 3-9.)
3-23

Figure 3-11.

Cargo Tie-Down Rings

3-52. SHOULDER HARNESS. Individual shoulder
harnesses may be installed at each seat. Each harness is connected to the upper fuselage structure and
to the seat safety belt buckle. Component parts should
be replaced as outlined in the preceding paragraph.
(Refer to figure 3-9.) Beginning with aircraft
U20602580, an inertia reel installation is offered.
Refer to figure 3-9 for installation.
3-53. GLIDER TOW-HOOK. A glider tow-hook,
which is mounted in place of tail tie-down ring, is
available for all models.
3-54. REAR VIEW MIRROR. A rear view mirror
may be installed on cowl deck above instrument panel.
Figure 3-11 shows details of rear view mirror installation.
3-55.

CARGO PACK.

3-56. REMOVAL.
a. Remove screws, fairing and seal from around
each landing gear spring.
b. Position a suitable support under pack.
c. Remove screws attaching pack to aircraft and
remove pack.
NOTE
If aircraft is to be returned to its original
configuration (minus cargo pack), the four
small panels which enclose area around
nose gear shock strut and drag brace may _
be left installed instead of the two larger
panels. However, the control extension
and cowl flap baffles must be removed as
outlined in paragraph 3-59.
3-57. INSTALLATION. Prior to positioning pack
under aircraft, inspect all rivnuts in bottom of fuselage for obstructions. Also check the small panels
which enclose area around nose gear shock strut
and drag brace. Two panels are provided in this
area on standard aircraft; these are to be replaced
by four smaller panels when a cargo pack is installed. If not previously removed, remove standard
panels by unsnapping quick-release fasteners. In3-24

Change 2

stall the smaller panels furnished with cargo pack.
NOTE
Install the rearmost panels first, right
hand panel lapping over left hand panel
along aircraft centerline. Install the
forward panels in a similar manner.
a. Move pack into position under aircraft. Raise
aft end of pack and place a support under it.
b. Raise forward end of pack and align two forward holes in pack rim with two front rivnuts. Install two screws to support forward end of pack.
NOTE
Install lock washers and flat washers under
heads of all pack attaching screws.
c. Raise aft end of pack and install two attaching
screws.
d. Check pack for proper alignment, install and
tighten all remaining screws, except for one screw
just forward and aft of each landing gear spring.
These two screws will be utilized later to help secure fairing which covers each landing gear opening.
e. Position rubber seal and fairing around each
main landing gear spring by spreading these components, at their split side, enough to slip them over
gear spring. When installed, split should be at back
of gear spring. Check alignment and proper fit of
fairing, then install fairing retaining screws.
NOTE
Seven screws are used to secure fairing
at each landing gear. Two screws, previously mentioned in step "d," secure top
of fairing and rim of cargo pack, in this
area, to fuselage. Five additional screws
secure and seal sides and bottom of each
fairing to pack.
f. Install cowl flap baffles and control extensions
in accordance with paragraph 3-60.
3-58. COWL FLAP BAFFLES AND CONTROL
EXTENSIONS. (Refer to figure 3-13.)

2

NOTE
Covers (1) and (3) are bonded to each other
around mirror (2) with a plastic bonding
agent, such as acetone.

/

Detail A
THRU 1971

1-

.

A
/-....

Detail A
1.
2.
3.
4.
5.
6.
7.

Cover
Mirror
Cover
Screw
Bracket
Washer
Knurled Nut

8. Cowl Deck
BEGINNING WITH 1972
9. Washer
10. Nut
11. Mirror Assembly
12. Spacer
13. Eyebrow
14. Washer
Figure 3-12.

.

.

.

Rear View Mirror Installation

3-59. REMOVAL.
a. Disconnect cowl flap control clevises (7) from
flaps and take off baffles (1) by removing screws (3)
and nuts (2).
b. Remove clevis (7) and link (5) from each control
end (8) and reinstall clevises,
c. Rig cowl flaps on standard aircraft per Section
12 and turbocharged aircraft per Section 12A.
3-60. INSTALLATION.
a. Disconnect cowl flap control clevises (7) from
flaps and remove clevises. Leave jam nuts (4) on
control ends (8).
b. Install links (5) on control ends (8), install jam
nuts (6) on links and attach clevises (7) to links. Do
not tighten jam nuts.
c. Position baffles (1) along sides of cowl flaps so
attaching holes are aligned and install attaching
screws and nuts.
NOTE
Each baffle is designed for installation on a
specific cowl flap. Determine correct baffle
for each flap. Turbocharged aircraft have
baffles as standard equipment. Note that
flanges on baffles are turned toward inside of
each cowl flap opening.
d. Check to ensure flexible controls reach their
internal stops in each direction. Mark controls so
full control travel can readily be checked and maintained during remaining rigging procedure.
e. Place cowl flap control lever in "OPEN" posttion and connect control ends (8) to flaps, but do not
secure at this time.
f. On standard aircraft, measure distance from
trailing edge of cowl skin. Disconnect clevises and
adjust links (5) and clevises (7) so each cowl flap

opens 6.00 inches with cockpit control OPEN and
1.05 inches with cockpit control CLOSED. On
turbocharged aircraft, adjust clevis to obtain measurements of 8.00 inches (cockpit control OPEN)
and 2. 50 inches (cockpit control CLOSED), then
secure clevises. These measurements are made in
a straight line from the aft edge of cowl flap, just
outboard of cutout to lower edge of firewall. Do not
measure from aft corners of cowl flap. If either
control needs to be lengthened or shortened, the
lower clamp may be loosened and housing slipped in
clamp or lower clevis may be adjusted. Maintain
sufficient thread engagement of clevis.
g. Check that locknuts are tight, clamps are secure, then cycle cowl flaps several times, checking
operation.
3-61.

CASKET CARRIER.

(Refer to figure 3-14.)

3-62. DESCRIPTION. An optional mortuary kit
consists of a casket carrier platform, rack assembly and belt tie-down assemblies. The kit provides aircraft modification instructions and parts
required to make the installation.
3-63. INSTALLATION. The following instructions
may be used to install platform, rack and tie-down
belts, and to load and secure casket:
a. Remove all seats and safety belts except pilot's
and copilot's.
b. Move pilot's and copilot's seats forward to their
limit of travel.
c. Attach belt assemblies to existing left forward
and left aft seat attach brackets as shown in detail
"G."

Change 3

3-25

STA 0 00

1/

9

FUSELAGE

2

H SIDE
STA
13.75

STA
34 50

STA
60 .00

LINE

DOOR

STA
84. 20

231/2

96 1/2
3
1. Baffles
2. Nut
3. Screw
4. Jam Nut
5. Link
6.

Jam Nut

7.
8.

Clevis
Control

2

4

COWL FLAP MODIFICATION

7

Figure 3-13.
3-26

Cargo Pack Installation

1.
2.
3.

Thumb Screw
Rack Assembly
Pad

4.
5.
6.
7.

Platform
Bracket
Seat Rail
Weld Assembly

8.
9.
10.

Washer
Nut
Bolt

3-27

UPPER-TO-FORWARD
BELT ATTACHMENT

UPPER BELT ATTACHMENT

LEFT SIDE FORWARD AND
AFT BELT ATTACHMENT

RIGHT SIDE AFT
BELT ATTACHMENT

RIGHT SIDE FORWARD
BELT ATTACHMENT

CARGO TIE-DOWN RING
LOWER BELT ATTACHMENT
INBOARD SEAT RAILS

Figure 3-14.

3-28

Casket Carrier Installation (Sheet 2 of 2)

d. Place platform in cabin and butt aft end of platform against step.
e. Secure both sides of platform to outboard seat
rails as shown in detail "A. "
f. Install rack on platform as shown in detail "B."
g. Install cargo tie-down rings on inboard seat
rails and attach lower belt as shown in detail "D."
NOTE
The cargo tie-down ring on left inboard
seat rail is tightened down against seat
rail, since no seat adjusting hole exists
in rail at this point. The cargo tiedown ring on right inboard seat rail will
engage an existing seat adjustment hole.

i. Attach upper belt to forward belt as shown in detail "C."
j. Attach right forward and right aft belts to existing seat belt attach points as shown in details "E"
and "F."
k. Remove pilot's seat back by removing quickrelease pins.
l. Load casket, adjusting end plates on rack
according to casket length. Tighten forward end
plate snugly.
m. Tighten all belts securely and recheck all tiedown attachments.
n. Reinstall pilot's seat back.
3-64. REMOVAL. After casket has been removed,
platform, rack, and belts may be removed by reversing installation procedure.

h. Attach upper belt at four points as shown in detail "H."

SHOP NOTES:

3-29/(3-30 blank)

SECTION 4
WINGS AND EMPENNAGE

TABLE OF CONTENTS
WINGS AND EMPENNAGE ..........
Wings
..............
Description
..............
Removal.
...........
Repair . . . . . . . . . ..
Installation.
...........
Adjustment
..............
Wing Struts .............
Description
.
.........
Removal and Installation ....

4-1.

WINGS AND EMPENNAGE.

4-2.

WINGS.

Page
...

4-1
4-1
4-1
. 4-1
4-2
. 4-2
4-2
. 4-2
4-2
4-2

(See figure 4-1.)

4-3. DESCRIPTION. Each all-metal wing panel is
a semicantilever, semimonocoque type, with two
main spars and suitable ribs for attachment of the
skin. Skin panels are riveted to ribs, spars and
stringers to complete the structure. Beginning with
U20601701 the leading edge skins are bonded. An
all-metal, balanced aileron, a flap, and a detachable
wing tip are mounted on each wing assembly. A
single rubberized bladder-type fuel cell is mounted
between the wing spars at the inboard end of each
wing and the leading edge of the left wing, thru 1971
models, has landing and taxi lights installed. Beginning with 1972 models the landing and taxi lights
are mounted in the lower engine nose cowl. Navigation/strobe lights are mounted at each contoured
wing tip.

Repair .. ....
.
.
4-2
Vertical Fin .
...........
.
. 4-2
Description .......
.......
4-2
Removal and Installation ........
4-2
Repair . . . . . .
.
4-2
Horizontal Stabilizer.
.........
.4-2
Description . .......
.....
. 4-3
Removal and Installation ....
. ..
4-3
Repair ..........
......
4-3

NOTE
To ease rerouting the cables, a guide wire
may be attached to each cable before it is
pulled free of the wing. Cable may then be
disconnected from wire. Leave guide wire
routed through the wing; it may be attached
again to the cable during reinstallation and
used to pull the cable into place.
f. Support wing at outboard end and disconnect strut
at wing fitting. Tie strut up with wire to prevent it
from swinging down and straining strut-to-fuselage
fittings. If the fuselage fitting projects from the
fuselage and is covered by the strut fairing, loosen
the fairing and slide it up the strut; the strut may
then be lowered without damage.
NOTE

4-4. REMOVAL. Wing panel removal is most easily
It is recommended that flap be secured in
accomplished if four men are available to handle the
streamlined position with tape during wing
wing. Otherwise, the wing should be supported with
removal to prevent damage, since flap will
a sling or maintenance stand when the fastenings are
swing freely.
loosened.
a. Remove wing gap fairings and screws securing
g. Mark position of wing attachment eccentric
cabin top skin to the wing top skin.
bushings (refer to figure 4-1); these bushings are
b. Remove all wing inspection plates.
used to rig out "wing-heaviness."
c. Drain fuel from cell of wing being removed.
h. Remove nuts, washers, bushings and bolts
d. Disconnect:
attaching wing spars to fuselage fittings.
1. Electrical wires at wing root disconnects.
2. Fuel lines at wing root. (Refer to preNOTE
cautions outlined in paragraph 13-3.)
3. Pitot line (left wing only) at wing root.
It may be necessary to rock the wing slightly
4. Cabin ventilator hose at wing root.
while pulling attaching bolts, or to use a long
e. Slack off tension on flap and aileron cables by
drift punch to drive out attaching bolts.
loosening turnbuckles, then disconnect cables at
flap and aileron bellcranks.
i. Remove wing and lay on padded stand.
Change 3

4-1

4-5. REPAIR. A damaged-wing panel may be repaired in accordance with instructions outlined in
Section 18. Extensive repairs of wing skin or structure are best accomplished using the wing repair jig,
which may be obtained from Cessna. The wing jig
serves not only as a holding fixture, making work on
the wing easier, but also assures absolute alignment
of the repaired wing.
4-6. INSTALLATION.
a. Hold wing in position and install bolts, bushings,
washers and nuts attaching wing spars to fuselage
fittings. Ensure eccentric bushings are positioned
as marked when removed.
b. Install bolts, spacers and nuts to secure upper
and lower ends of wing strut to wing and fuselage
fittings.
c. Route flap and aileron cables, using guide wires.
(See note in paragraph 4-4. )
d. Connect:
1. Electrical wires at wing root disconnects.
2. Fuel lines at wing root. (Refer to precautions outlined in paragraph 13-3.)
3. Pitot line (if left wing is being installed. )
4. Wing leveler vacuum line, if installed, at
wing root.
5. Ventilator hose at wing root.
e. Rig aileron system (Section 6).
f. Rig flap system (Section 7).
g. Refuel fuel cell and check for leaks.
h. Check operation of navigation/strobe also landing and taxi lights thru 1971 models.
i. Check operation of fuel quantity indicator.
j. Install wing gap fairings.
NOTE
Be sure to insert soundproofing panel in
wing gap, if such a panel was installed
originally, before replacing wing root
fairings.
k. Install all wing inspection plates, interior panels
and upholstery.
1. Test operate flap and aileron systems.
4-7. ADJUSTMENT (CORRECTING "WING-HEAVY"
CONDITION). If considerable control wheel pressure
is required to keep the wings level in normal flight,
a "wing-heavy" condition exists.
a. Remove wing fairing strip on "wing-heavy" side
of aircraft.
b. (See figure 4-1. ) Loosen nut (7) and rotate bushings (5) simultaneously until the bushings are positioned with the thick side of the eccentrics up. This
will lower the trailing edge of the wing, and decrease
"wing-heaviness" by increasing the angle-of-incidence
of the wing.
CAUTION
Be sure to rotate the eccentric bushings
simultaneously. Rotating them separately
will destroy the alignment between the offcenter bolt holes in the bushings, thus exerting a shearing force on the bolt, with possible damage to the hole in the wing spar.
4-2

Change 3

c. Tighten nut and reinstall fairing strip.
d. Test-fly the aircraft. If the "wing-heavy" condition still exists, remove fairing strip on the "lighter"
wing, loosen nut and rotate bushings simultaneously
until the bushings are positioned with the thick side
of the eccentric down. This will raise the trailing
edge of the wing, thus increasing "wing heaviness"
to balance heaviness in the opposite wing.
e. Tighten nut, install fairing strip and repeat
flight test.
4-8.

WING STRUTS.

(See figure 4-2.)

4-9. DESCRIPTION. Each wing has a single lift
strut which transmits a part of the wing load to the
lower portion of the fuselage. The strut consists of
a streamlined tube riveted to two end fittings for
attachment at the fuselage and wing.
4-10. REMOVAL AND INSTALLATION.
a. Thru U20602501 remove screws from strut
fairings and slide fairing along strut. Beginning
with U20602501 the upper strut fairing is split along
the aft edge and attached together with screws for
easy removal.
b. Remove fuselage and wing inspection plates at
strut junction points.
c. Support wing securely, then remove nut and
bolt securing strut to fuselage.
d. Remove nut, bolt and spacer used to attach
strut to wing, then remove strut from aircraft.
e. Reverse preceding steps to install strut.
4-11. REPAIR. Wing strut repair is limited to replacement of tie-downs and attaching parts. A badly
dented, cracked or deformed wing strut must be replaced.
4-12.

VERTICAL FIN.

(See figure 4-3.)

4-13. DESCRIPTION. The fin is primarily of metal
construction, consisting of ribs and spars covered
with skin. Fin tips are of glass fiber of ABS construction. Hinge brackets at the rear spar attach
the rudder.
4-14. REMOVAL AND INSTALLATION. A fin may
be removed without first removing the rudder. However, for access and ease of handling, the rudder
may be removed by following procedures outlined in
Section 10.
a. Remove fairings on either side of fin.
b. Disconnect flashing beacon lead, tail navigation
light lead, antennas and antenna leads, and rudder
cables, if rudder has not been removed.
c. Remove screws attaching dorsal to fuselage.
d. Remove bolts attaching fin front and rear spars
to fuselage, and remove vertical fin.
e. Install fin by reversing preceding steps. Be
sure to check and reset rudder and elevator travel
if any stop bolts were removed or settings disturbed.
4-15. REPAIR. Fin repair should be accomplished
in accordance with applicable instructions outlined
in Section 18.
4-16.

HORIZONTAL STABILIZER (See figure 4-4.)

4-17. DESCRIPTION. The horizontal stabilizer is
primarily of metal construction, consisting of ribs
and a front and rear spar which extend throughout
the full spars and ribs. Stabilizer tips are of ABS
construction. The elevator tab actuator screw is
contained within the horizontal stabilizer assembly,
and is supported by a bracket riveted to the rear
spar. The underside of the stabilizer contains a
covered opening which provides access to the elevator tab actuator screw. Hinge brackets at the rear
spar support the elevators.
4-18. REMOVAL AND INSTALLATION.
a. Remove elevators and rudder in accordance with
procedures outlined in Sections 8 and 10.
b. Remove vertical fin in accordance with proce-

dures outlined in paragraph 4-14.
c. Disconnect elevator trim control cables at
clevis and turnbuckle inside tailcone, remove pulleys
which route aft cables into horizontal stabilizer, and
pull cables out of tailcone.
d. Remove bolts securing horizontal stabilizer to
fuselage.
e. Remove horizontal stabilizer.
f. Install horizontal stabilizer by reversing preceding steps. Rig control systems as necessary.
Check operation of tail navigation light and flashing
beacon.
4-19. REPAIR. Horizontal stabilizer repair should
be accomplished in accordance with applicable procedures outlined in Section 18.

SHOP NOTES:

Change 1

4-3

2

2

Detail

B

Detail A

* NOTE
9*

Coat bolt and hole with
Electro Moly No. (MILG-21164) grease.

NOTE
*The forward bushing is approximately
half the length of the aft bushing. Care
should be taken to install the short bushing in the forward side and the long
bushing in the aft side.

13
11

** Beginning with serials

U20603021 wing
fuel bay cover panels are of bonded construction.

1.
2.
3.
4.
5.
6.
7.

Nut
Washer
Bolt
Bolt
Bushing
Washer
Nut

8.
9.
10.
11.
12.
13.
14.

Change 3

*THRU AIRCRAFT SERIAL
U20601700

Tip Assembly
Landing and Taxi Light
Fuel Filler Cap
Fillet
Fairing
Flap
Aileron
Figure 4-1.

4-4

12

Wing Installation

* NOTE
Beginning with aircraft serial
U20602502 wrap strut using
Y8562 polyurethane tape (1"
wide) centered at point where
strut cuff terminates.

4

9

7

10

1

9
4

7

14

*

* NOTE

*15

*THRU

U20601700

* BEGINNING WITH U20601701

1.
2.
3.
4.
5.

Screw
Upper Fairing
Bolt
Washer
Cotter Pin

6.
7.
8.
9.
10.

Rivet
Strut Fitting
Pin
Nut
Spacer

Coat bolt and hole with
Electro Moly No. (MILG-21164) grease.

11. Mooring Ring
12. Spring
13. Fuselage Fitting
14. Lower Fairing
15. Tape

Figure 4-2. Wing Strut Installation
Change 3

4-5

NOTE
On Aircraft Serials U20601595 thru
U20601618, and U20601633 & On,
center hole in aft fin attach fitting
(5) has been drilled to accept AN6
Bolts. On Aircraft Serials U20601619
thru U20601632, center hole will
accept AN5 Bolt.

B

Detail

Detail A

B

THRU U20601905
D
DetailD

/

D

10

12Detail
l

\
* THRU U20601905
Detail C

/Jv L 1-'* \

56

\]

/ THRU AIRCRAFT SERIAL P20600648
and U20601587 when not modified per
Single-engine Service Letter SE71-29,
Dated October 15. 1971, use washers
( 7) and (8) on rear fin fitting. Use
\
a10 washers (7). (8) and (9) when modified
by installation of new bulkhead. and all
ii11
Service Parts. Use washers (7) when
modified by reaming of bolt holes.

BEGINNING WITH U20601906
Detail D
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.

Fin Assembly
Upper Rudder Hinge
Center Rudder Hinge
Lower Rudder Hinge
Aft Attach Fitting
Bolt
Washer
Washer
Washer
Nut
Washer
Fwd Attach Fitting

Change 3

/

/

H

*

AIRCRAFT SERIALS U20601588 THRU
U20601904, use washers (7), (8) and (9).

TORQUE AN5 BOLTS TO 140-225 LB IN.
TORQUE AN6 BOLTS TO 190-390 LB IN.
TORQUE AN7 BOLTS TO 500-840 LB IN.
NOTE
Beginning with 1962 Models, Cessna
Single-engine Service Letter SE72-3
dated, February 11, 1972 should be
complied with.
Figure 4-3.

4-6

/

Vertical Fin Installation

BEGINNING WITH U20601906
Detail C

2

3

6
6

D

I *14

*
4>

/

<

* NOTE

1. Nutplate
2. Washer
3. Bolt
4. Bracket
5. Nut
6. Washer

i.DetailDetail D
D

Bo,-ilii

An abrasion boot kit may be
obtained from the Cessna
Service Parts Center.

7.
8.
9.
10.
11.
12.

Bracket
Bolt
Elevator Pylon Bracket
Elevator Inboard Hinge
Elevator Outboard Hinge
Upper Right Fairing

Figure 4-4.

13.
14.
15.
16.
17.
18.

Upper Left Fairing
Abrasion Boot
Lower Left Moulding
Lower Right Moulding
Forward Left Fairing
Forward Right Fairing

Horizontal Stabilizer
Change 1

4-7/(4-8 blank)

SECTION 5

TABLE OF CONTENTS

LANDING GEAR AND BRAKES
Page

5-1
....
.
. .....
LANDING GEAR
5-1
...........
Description
5-2
Main Landing Gear ...........
. 5-2
...
...
Trouble Shooting ....
..
.. 5-4A
..
.
Removal. ..
. 5-4A
Installation ..........
Removal and Installation of Main
Landing Gear Brake Fairings . . .5-4A
Removal and Installation of Standard
. 5-4A
Main Wheel Speed Fairings ...
Main Wheel and Tire Assembly . . . 5-4A
5-4A
Description ..........
Removal of Main Wheel and
..
5-4A
Tire Assembly .......
Disassembly of Cleveland Main
Wheel and Tire Assembly . . .5-4A
Inspection and Repair of
Cleveland Main Wheel and
5-4B
Tire Assembly .......
Reassembly of Cleveland Main
Wheel and Tire Assembly . . . 5-4B
Disassembly of McCauley Main
. 5-4B
Wheel and Tire Assembly .
Inspection and Repair of
McCauley Main Wheel and
. 5-4B
Tire Assembly ........
Reassembly of McCauley Main
Wheel and Tire Assembly .. .5-4B
Main and Nose Wheel ThruBolt Nut or Capscrew Torque
Values ...........
5-4C
Installation of Main Wheel and
5-4C
Tire Assembly ........
.
5-5
Removal of Main Wheel and Axle .
5-5
Installation of Main Wheel and Axle. .
5-5
Main Wheel Alignment .......
5-5
Wheel Balancing ..........
..
5-5
Step Bracket Installation ...
Brake Line Fairing Replacement . . . 5-8
5-8
Nose Gear .
....................
Trouble Shooting ..........
5-8
Replacement of Nose Gear. ......
5-9
Standard Nose Gear Speed Fairing
Replacement ...........
5-9
Heavy-Duty Nose Wheel Speed
5-11
Fairing Adjustment ........
Nose Wheel and Tire Assembly . .. 5-11
Description
.
. . . ..
. 5-11

5-1.

LANDING GEAR.

5-2. DESCRIPTION. These aircraft are equipped
with non-retractable, tricycle landing gear, utilizing
flat spring-steel main gear struts. Disc-type brakes
and tube-type tires are installed on the axle at the
lower end of the strut. Speed fairings or heavy-duty
wheels may be installed on some aircraft. The nose
gear is a combination of a conventional air/oil (oleo)

Removal of Nose Wheel and
..
Tire Assembly ..
Disassembly of Cleveland Nose
· .
Wheel and Tire Assembly
Inspection and Repair of
Cleveland Nose Wheel and
Tire Assembly ........ .
Reassembly of Cleveland Nose
.
Wheel and Tire Assembly .
Disassembly of McCauley Nose
Wheel and Tire Assembly . .
Inspection and Repair of
McCauley Nose Wheel and
.
......
Tire Assembly .
Reassembly of McCauley Nose
Wheel and Tire Assembly · .
Installation of Nose Wheel and
......
Tire Assembly .
Standard Nose Gear Strut ......
Description ........
.
Disassembly ..........
Reassembly ..........
Heavy-Duty Nose Gear Strut . ....
Description .
...
...
.
........
Disassembly .
.
.....
Reassembly .
Wheel Balancing .
.........
Torque Links .
..........
..
.
Shimmy Dampener .....
.....
Nose Wheel Steering System
Description
..
........
.
Removal and Installation ....
.
Rigging ..........
..... ..
.
Brake System ......
.
......
Description . ...
......
Trouble Shooting .
. .
Brake Master Cylinders. ...
Removal and nstallation . . ..
Disassembly and Repair ..
Hydraulic Brake Lines .......
Wheel Brake Assemblies ......
Removal.
.... ...
Inspection and Repair ......
Assembly
..........
Installation ..........
Checking Brake Lining Thickness .
Brake Lining Replacement. .....
Brake Bleeding.
.
.........
Parking Brake System .....

5-11
5-11
5-11
5-11
5-11

5-12
5-12
5-12
5-12
5-12
5-12
5-14
5-14
5-14
5-14
5-16
5-17
5-17
5-17
5-17
5-17
5-17
5-19
5-19
5-19
5-19
5-21
5-21
5-21
5-21
5-21
5-21
5-21
5-21
5-21
5-21
5-21
5-24
5-24

strut and fork, incorporating a shimmy dampener.
The nose wheel is steerable with the rudder pedals
up to a maximum pedal deflection, after which it
becomes free-swiveling up to a maximum travel right
or left of center. Through the use of the brakes, the
aircraft can be pivoted around the outer wing strut
fitting. A speed fairing or a heavy-duty shock strut
and wheel may be installed on some aircraft.
Change 3

5-1

5-3.

MAIN LANDING GEAR.

5-4.

TROUBLE SHOOTING.
TROUBLE

AIRCRAFT LEANS TO
ONE SIDE.

UNEVEN OR EXCESSIVE
TIRE WEAR.

WHEEL BOUNCE EVIDENT
EVEN ON SMOOTH SURFACE.

SHOP NOTES:

5-2

PROBABLE CAUSE

REMEDY

Incorrect tire inflation.

Inflate to correct pressure.

Landing gear attaching
parts not tight.

Tighten loose parts; replace
defective parts.

Sprung landing gear spring.

Replace spring.

Bent axle.

Replace axle.

Different quantity of fuel
in wing cells.

Refuel aircraft.

Structural damage to landing
gear bulkhead components.

Replace damaged parts.

Incorrect tire inflation.

Inflate to correct pressure.

Wheels out of alignment.

Align wheels. See figure 5-2.

Wheels out of balance.

Refer to paragraph 5-16.

Sprung landing gear spring.

Replace spring.

Bent axle.

Replace axle.

Dragging brake.

Refer to paragraph 5-48.

Wheel bearings not adjusted
properly.

Tighten axle nut properly.

Out of balance condition.

Correct in accordance with 5-16.

~~~~*
,„,
Step assembly (9)
right
not used
used on
on right
gear strut of air- *30
craft equipped with
\
cargo doors.

= -^not
2 I~J~
I
- ^iJ /,'1
^...-' /
|I
2
.^
.I .
-

33

fe

*y

(HEAVY-DUTY)
THRU U20603020

o0 

l1

VY

24
i

of bolts (28).

1.

2.
3.
4.
5.
6.
7.
8.
9.
10.

Screw

Lockwasher
Grease Seal Ring
Felt Seal
Grease Seal Ring
Bearing Cone
Outer Wheel Hall
Tire and Tube
Inner Wheel Half
Grease Seal Plate and Felt

Figure 5-11.
5-22

Change I

I

/

a

11.

Screw

12.
13.
14.
15.
16.
17.
18.
19.
20.

Washer
Grease Seal Ring
Bearing Cup
Brake Disc
Torque Plate
Pressure Plate
Anchor Bolt
Hydraulic Fitting
Washer

Wheels and Brakes (Sheet 1 of 2)

21.

Nut

22.
23.
24.
25.
26.
27.
28.
29.
30.

Bolt
Washer
Bleeder Cap
Brake Cylinder
Piston
Brake Lining
Bolt
Brake Lining
Back Plate

3

STEEL FLANGE

2
1

I
Washer

13

:/

/'

// /

>

12

)8

STEEL FLANGE

McCAULEY WHEEL AND BRAKE

17

1. Washer
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.

Cap Screw
Retainer Ring
Grease Seal Retainer
Grease Seal Felt
Grease Seal Retainer
Bearing Cone/
Wheel Flange
Spacer
Tire
Tube
Hub Assembly
Lining
Back Plate Assembly
Disc Assembly
Torque Plate
Pressure Plate
Guide Pin
Bleeder Fitting
Bleeder Screw
Dust Cap
Cylinder
O-Ring
Piston
Lining
Thru-Bolt-

Figure 5-11.

la
i<

(4

\'

-

r
2

/
1

/

/
/

/
1

19

,

2

5
//

*$

/
/20

24

r

23
2

Wheels and Brakes (Sheet 2 of 2)
Change 3

5-22A

MAIN

NOSE

GEAR

GEAR

TIRE
WHEEL NUMBER

NUT/CAPSCREW WHEELHALF

SIZE

MANUFACTURER

FLANGE

TORQUE

x

C163001-0301

6.00X6

CLEVELAND

150 Ib-in.

MAGNESIUM

X

C163001-0302

8 .00X6

CLEVELAND

150 Ib-in.

MAGNESIUM

X

C163002-0103

6.00X6

McCAULEY

90-100 Ib-in.

ALUMINUM

00X6

McCAULEY

90-100 lb-in.

ALUMINUM

C163004-0102

6. 00X6

McCAULEY

C163004-0101

8.00X6

McCAULEY

190-200 Ib-in.

ALUMINUM

1241156-12

5.00X5

CLEVELAND

90 lb-in.

MAGNESIUM

1241156-11

6.00X6

CLEVELAND

150 lb-in.

MAGNESIUM

X

C163002-0201

5.00X5

McCAULEY

90-100 Ib-in.

X

C 163003-0201

5.00X5

McCAULEY

90-100 Ib-in.

STEEL

X

C163003-0301

6.00X6

McCAULEY

*190-200 lb-in.

STEEL

X

C163003-0401

5.00X5

McCAULEY

'190-200 Ib-in

STEEL

I

X
X

I

X
X
X

i

C163002-0104

Figure 5-11A.

|8

1190-200

Ib-in.

1

ALUMINUM

ALUMINUM

Landing Gear Wheel Thru-Bolt Nut and Capscrew Torque Values

-Capscrews

SHOP NOTES:

5-22B

Change 3

D2007C3-13 Temporary Change 2
22 February 1978

13. Spring

.

Cable Assembly

20.

Bracket20601668

23. Gasket
1916

2f

"1

t=I

5-23

D

Detail D
1.

Attaching Angle
2. Stiffener Angle
3. Parking Brake Handle
4. Housing
5. Clamp
6.
Cotter Pin
7. Positioning Pin
8.
Cable Assembly

9.
10.
11.
12.
13.
14.
15.

Bracket
Bellcrank
Cable
Pin
Spring
Spacer
Pulley

16.
17.
18.
19.
20.
21.
22.
23.

Brake Line
Clamp
Brake Master Cylinder
Brake Hose
Bracket
Elbow
Nut
Gasket

Figure 5-12.

Brake System
5-23

rivet. While holding the back plate down firmly
against the lining, hit the punch with a hammer to
set the rivet. Repeat blows on the punch until lining
is firmly against the back plate.
g. Realign the lining on the back plate and install
rivets in remaining holes.
h. Install a new lining on pressure plate in the same
manner.
i. Position pressure plate on anchor bolts, and
place cylinder in position so the anchor bolts slide
into torque plate.
j. Install the back plates with bolts and washers.
Safety wire the bolts.
5-60. BRAKE BLEEDING. Standard bleeding, with
a clean hydraulic pressure source connected to the
wheel cylinder bleeder, is recommended.
a. Remove brake master cylinder filler plug and
screw a flexible hose with a suitable fitting into the
filler hole. Immerse the free end of the hose in a
container with enough hydraulic fluid to cover the
end of the hose.
b. Connect a clean hydraulic pressure source, such
as a hydraulic hand pump or Hydro Fill unit, to the
bleeder valve in the wheel cylinder.
c. As fluid is pumped into the system, observe the

SHOP NOTES:

5-24

immersed end of the hose at the brake master cylinder for evidence of bubbles being forced from the
brake system. When bubbling has ceased, remove
the bleeder source from the brake wheel cylinder and
tighten the bleeder valve.
NOTE
Ensure that the free end of the hose from the
brake master cylinder remains immersed
during the entire bleeding process.
d. Remove hose from brake master cylinder and
replace filler plug. Be sure vent hole in filler plug
is open.
5-61.

PARKING BRAKE SYSTEM.

5-62. DESCRIPTION. The parking brake system is
essentially a ratchet-held handle which depresses
and holds the brake master cylinders in the compressed position. No adjustment is provided in the system.
Replacement of worn or defective parts will restore
the system to its correct operation. Figure 5-12 may
be used as a guide for replacement of parts.

SECTION 6
AILERON CONTROL SYSTEM

TABLE OF CONTENTS

Page

AILERON CONTROL SYSTEM .........
Description ............
Trouble Shooting ............
Control Column .............
Description ................
Removal and Installation ......
.........
Repair.
Bearing Roller Adjustment ......
Aileron Bellcrank .........
..

6-1
6-1
6-1
6-2
6-2
. 6-2
6-6
6-7
. 6-7

6-1. AILERON CONTROL SYSTEM.
ure 6-1.)
6-2.

DESCRIPTION.

6-3.

TROUBLE SHOOTING.

Removal and Installation
Repair ...........
Cables and Pulleys ............
Removal and Installation
..............
Ailerons.
Removal and Installation
...............
Repair
Rigging ............

(Refer to fig-

.......
.......
.....
....

6-7
6-7
6-8
. 6-8
6-8
. 6-8
6-8
6-8

comprised of push-pull rods, bellcranks, cables,
pulleys, quadrants and components forward of the
instrument panel, all of which, link the control
wheels to the ailerons.

The aileron control system is

NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 6-17.

TROUBLE
LOST MOTION IN CONTROL
WHEEL.

RESISTANCE TO CONTROL
WHEEL MOVEMENT.

PROBABLE CAUSE

REMEDY

Loose control cables.

Check cable tension. Adjust
cables to proper tension.

Broken pulley or bracket,
cable off pulley or worn
rod end bearings,

Check visually. Replace worn or
broken parts, install cables
correctly.

Cables too tight.

Check cable tension. Adjust
cables to proper tension.

Pulleys binding or cable off.

Observe motion of the pulleys.
Check cables visually. Replace
defective pulleys. Install cables
correctly.

Bellcrank distorted or
damaged.

Check visually.
bellcrank.

Replace defective

Defective quadrant assembly.

Check visually.
quadrant.

Replace defective

Clevis bolts in system too
tight.

Check connections where used.
Loosen, then tighten properly
and safety.
Change 3

6-1

6-3.

TROUBLE SHOOTING (Cont).
TROUBLE

CONTROL WHEELS NOT
LEVEL WITH AILERONS
NEUTRAL.

PROBABLE CAUSE

REMEDY

Improper adjustment of
cables.

Refer to paragraph 6-17.

Improper adjustment of
aileron push-pull rods.

Adjust push-pull rods to obtain
proper alignment.

DUAL CONTROL WHEELS
NOT COORDINATED.

Cables improperly adjusted.

Refer to paragraph 6-17.

INCORRECT AILERON
TRAVEL.

Push-pull rods not adjusted
properly.

Refer to paragraph 6-17.

Incorrect adjustment of travel
stop bolts.

Refer to paragraph 6-17.

6-4.

CONTROL COLUMN.

(Refer to figure 6-2.)

6-5. DESCRIPTION. Rotation of the control wheel
rotates four bearing roller assemblies (3) on the end
of the control wheel tube (4), which in turn, rotates
a square control tube assembly (18) inside and extending from the control wheel tube (4). Attached to
this square tube (18) is a quadrant (32) which operates the aileron system. This same arrangement is
provided for both control wheels. Synchronization of
the control wheels is obtained by the interconnect
cable (38), turnbuckle (37) and adjustment terminals
(35). The forward end of the square control tube (18)
is mounted in a bearing block (27) on firewall (33)
and does not move fore-and-aft, but rotates with the
control wheel. The four bearing roller assemblies
(3) on the end of the control wheel tube reduce friction as the control wheel is moved fore-and-aft for
elevator system operation. A sleeve weld assembly
(7), containing bearings which permit the control
wheel tube to rotate within it, is secured to the control wheel tube by a sleeve and retaining ring in such
a manner it moves fore-and-aft with the control wheel
tube. This movement allows the push-pull tube (19)
attached to the sleeve weld assembly (7) to operate an
elevator arm assembly (22), to which one elevator
cable (39) is attached. A torque tube (21) connects
this arm assembly (22) to the one on the opposite end
of the torque tube (21), to which the other elevator
cable is attached. When dual controls are installed,
the copilot's control wheel is linked to the aileron and
elevator control systems in the same manner as the
pilot's control wheel.
6-6. REMOVAL AND INSTALLATION.
a. THRU AIRCRAFT SERIAL 20601700. (Refer to
figure 6-2, sheet 1.) Remove screws attaching control wheel (2) to control wheel tube assembly (4) and
remove wheel. Disconnect electrical wiring to map
light and mike switch, if installed.

6-2

b. BEGINNING WITH AIRCRAFT SERIAL 20601701. (Refer to figure 6-2, sheet 2.) Slide cover
(2) toward instrument panel to expose adapter (3).
Remove screws securing adapter (3) to control wheel
tube assembly (1) and remove control wheel assembly. Disconnect electrical wiring to map light, mike
switch and electric trim switch at connector (18), if
installed. Slide cover (2) off control wheel tube
assembly (1).
c. (Refer to figure 6-2, sheet 1.) Remove decorative cover from instrument panel.
d. Remove screw securing adjustable glide plug
(16) to control tube assembly (18) and remove plug
(16) and glide (17).
e. Disconnect push-pull tube (19) at sleeve weld
assembly (7).
f. THRU AIRCRAFT SERIAL 20601700. (Refer to
figure 6-2, sheet 1.) Remove screws securing cover
plate (15 or 24) at instrument panel.
g. BEGINNING WITH AIRCRAFT SERIAL 20601701. (Refer to figure 6-2, sheet 2.) Remove
screws securing cover plate (20) at instrument panel.
h. (Refer to figure 6-2, sheet 1.) Using care, pull
control wheel tube assembly (4) aft and work assembly out through instrument panel.
NOTE
To ease removal of control wheel tube assembly (4), snap ring (11) may be removed from
its locking groove to allow sleeve weld assembly (7) additional movement.
If removal of control tube assembly (18) or
quadrant (32) is necessary, proceed to step
"i."
i. Remove safety wire and relieve direct cable
tension at turnbuckles (index 9, figure 6-1).

1

2
Detail

DetalA

3

REFER TO FIGURE 6-3

DetailH
REFER TO
FIGURE 6-3
3

Detail G

NOTE
Detail

F

Shaded pulleys used
in this system.

REFER TO
FIGURE 6-2

CAUTION
1. Pulley

2.
3.
4.

Spacer
Cable
Guard
Spar
Rear Carry-Thru

(Outbrd Drect)
Wing Spar
?.. Cable
8. Cable (Outboard Direct)

9.

Turnbuckle (Direct Cable)

14.
10.
11.
12.

(Interconnect
Turnbuckle
Cable
(Carry-Thru)
rect)NON.
(Inbod
CableStrip
Rub

14.

(Interconnect
Turnbuckle
Cable)
Cable)

Figure 6-1.

MAINT AIN PROPE R CON TROL

40 LBS
M

ERON CARRY10 LBS ON AIPROP

CARRY± 10 LBS ON AILERON
40 LBS
MPENSION:
TE
CABLE
THRU
THRU CABLE (AT AVERAGE TEMPERATURE FOR THE AREA.)
REFER TO FIGURE 1-1 FOR TRAVEL.

Aileron Control System
Change 1

6-3

1

2

LEFT HAND
/CONTROL
COLUMN

Allow 0.030 " maximum clearance
between bearing block (27) and nut
(34) after tightening.

9
9

Adjust interconnect cable (38) tension to 30 ± 10 lbs.

4

*Safety wire these items.

12

24

12

A

/

-

23

Detail A
THRU AIRCRAFT
SERIAL 20601700

RIGHT HAND
CONTROL COLUMN

/

2.
3.

.

4.
5.
6.
7.

25

11

-

--

22*

19

*

PER
SIDE

VIEW A-A
* NOTE
Refer to Section 8
32 27-

.

38

*35

*36

NOTES
Washers (26) are of various thicknesses
and are used to obtain dimension shown
in view A-A.

27
34
/

Use washers (6) as required to adjust free
play. Do not exceed 4 washers per assembly.
Used only on aircraft NOT equipped with dual
control wheel installation.
Figure 6-2.
6-4

Change 2

Collar
1.
Decorative
Control Wheel
Bearing Roller
Assembly
Control Wheel Tube
Assembly
Collar
Washer
Sleeve Weld Assembly

8.
9.
10.
11.
12.

Bearing
Bearing Race
Thrust Bearing
Snap Ring
Grommet

14.

Spacer
Collar

15.
16.
1 .
18.
19.
20.
21.
.22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.

*35.
36.
37.

B
38.

Control Column Installation (Sheet 1 of 2)

Cover Plate
Adjustable Glide Plug
Control Tube Glide
Control Tube Assembly
Push-Pull Tube
Guide Assembly
Deleted
Arm Assembly (Elev)
Retainer
Cover Plate
Retainer
Washer
Bearing Block
Support Bracket
Idler Shaft
Cable Guard
Cable (Aileron Direct)
Quadrant (Cable Drum)
Firewall
Nut
Adjustable Nut
Roll Pin
Turnbuckle (Interconnect Cable)
Cable (Interconnect)

1.
2.
3.
4.
5.
6.
7.
8.
9.

Tube Assembly
Cover
Adapter
Rubber Cover
Plate
Map Light Rheostat
Terminal Block
Map Light Assembly
Control Wheel

10.
11.
12.
13.
14.
15.
16.
17
18.
Figure 6-2.

Pad
Mike Switch
Plug
Insulator
Electric Trim Switch
Plug
Bracket
Cable
Connector

19.
20.
21.
22.
23.
24.
25.
26.
27.

Bearing Assembly
Cover Plate
Collar
Single Controls Cover Plate
Bearing Block
Firewall
Guard Assembly
Quadrant
Cover Strap

Control Column Installation (Sheet 2 of 2)
Change 3

6-5

12

14

3

X\

"V
"'

19

>

y^^"NOTE
Push-pull tube (9) is adjustable
at the aileron end only.

tJL--"--24
1is-_____

0

21

/

'
22/.^~
~threads

1. Hinge Support
2. Aileron Assembly
3. Balance Weight
4. Upper Skin
5. Bolt
6. Leading Edge Skin
7. Plug Button
8. Access Plate
9. Push-Pull Tube

Prior to rigging the aileron sys-

tem, remove stop bolts (11) and
apply locking compound, LOCTITE
GRADE C OR EQUIVALENT, to
and reinstall in bellcrank.

10. Retainer
11. Travel Stop Bolt
12. Pivot Bolt
13. Upper Wing Skin
14. Upper Aileron Bellcrank Bracket
15. Bushing
16. Bearing
17. Cable (Outboard Direct)
18. Cable (Carry-Thru)
Figure 6-3.

j. Remove safety wire, relieve interconnect cable
tension at turnbuckle (37) and remove cables from
quadrant (32).
k. Remove safety wire and remove roll pin (36)
through quadrant (32) and control tube assembly (18).
1. Remove pin, nut (34) and washer from control
tube assembly (18) protruding through bearing block
(27) on forward side of firewall (33).
m. Using care, pull control tube assembly (18) aft
and remove quadrant (32).
n. Reverse the preceding steps for reinstallation.
Rig aileron and elevator control systems in accor-

6-6

Change 1

7

19.
20.
21.
22.
23.
24.
25.
26.

Bellcrank
Cable Guard
Lower Aileron Bellerank Bracket
Lower Wing Skin
Brass Washer
Cable Lock
Nutplate
Bolt

Aileron Installation

dance with paragraphs 6-17 and 8-13 respectively.
Safety turnbuckles and all other items previously
safetied. Tighten nut (34) securing control tube
assembly (18) to firewall snugly, then loosen nut to
0. 030" maximum clearance between nut and bearing
block, align cotter pin hole and install pin.
6-7. REPAIR. Worn, damaged or defective shafts,
bearings, quadrants, cables or other components
should be replaced. Refer to Section 2 for lubrication requirements.

AVAILABLE FROM CESSNA SERVICE PARTS CENTER (TOOL NO. SE 716)

Figure 6-4.

Inclinometer for Measuring Control Surface Travel

6-8. BEARING ROLLER ADJUSTMENT. (BEGINNING WITH AIRCRAFT SERIAL 20601701.) (Refer
to figure 6-2.) Each bearing assembly (index 19,
sheet 2) has an 0.062" eccentric adjustment when
installed, for aligning the control tube weld assembly
(index 7, sheet 1) and push-pull tube (index 19, sheet
1) with the guide assembly (index 20, sheet 1). For
alignment, proceed as follows:
a. Remove control wheel assembly in accordance
with paragraph 6-6.
b. Install cover plate (index 20, sheet 2) backwards
(bearings on aft side) and leave loose with instrument
panel,
c. Align control wheel tube assembly (index 4, sheet
1) for free travel of push-pull tube (index 19, sheet 1)
along full length of guide assembly (index 20, sheet
1).
d. Center cover plate (index 20, sheet 2) over tube
and bearing assembly and secure plate to instrument
panel.
e. Adjust each bearing (index 19, sheet 2) to control
wheel tube assembly and tighten bearings in place.
f. Remove cover plate and reinstall with bearings
facing forward.
6-9.

AILERON BELLCRANK.

(Refer to figure 6-3.)

b. Remove safety wire and relieve cable tension at
turnbuckles (index 9, figure 6-1).
c. Disconnect control cables from bellcrank (19).
d. Disconnect push-pull tube (9) at bellcrank (19).
e. Remove bolt securing bellcrank to wing structure.
f. Remove bellcrank through access opening. using
care that bushing (15) is not dropped from bellcrank.
NOTE
Brass washers (23) may be used as shims
between lower end of bellcrank and lower
bracket (21). Retain these shims. Tape
open ends of bellcrank to prevent dust and
dirt from entering bellcrank needle bearings (16).
g. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 6-17,
safety turnbuckles and reinstall all items removed
for access.
6-11. REPAIR. Repair of bellcranks consists of
replacement of defective parts. If needle bearings
are dirty or in need of lubrication, clean thoroughly
and lubricate as outlined in Section 2.

6- 10. REMOVAL AND INSTALLATION.
a. Remove access plate inboard of each bellcrank
(19) on underside of wing.
6-7

6-12.
6-1.)

CABLES AND PULLEYS.

(Refer to figure

6-13. REMOVAL AND INSTALLATION.
a. Remove access plates, wing root fairings and
upholstery as required.
b. Remove safety wire and relieve cable tension at
turnbuckles (9 and 13).
c. Disconnect cables from aileron bellcranks (index
19, figure 6-3) and quadrants (index 32, figure 6-2).
d. Remove cable guards and pulleys as necessary
to work cables free of aircraft.
NOTE
To ease routing of cables, a length of wire
may be attached to end of cable before
being withdrawn from aircraft. Leave
wire in place, routed through structure;
then attach cable being installed and use
to pull cable into position.
e. Reverse the preceding steps for reinstallation.
f. After cables are routed in position, install pulleys and cable guards. Ensure cables are positioned
in pulley grooves before installing guards.
g. Re-rig aileron system in accordance with paragraph 6- 17, safety turnbuckles and install access
plates, fairings and upholstery removed in step "a."
6-14.

AILERONS.

(Refer to figure 6-3. )

6-15. REMOVAL AND INSTALLATION.
a. Remove access plates (8) and plug buttons (7)
from underside of aileron.
b. Disconnect push-pull tube (9) at ailerons.
c. Remove bolts (5) attaching ailerons to hinge
supports (1).
d. Using care pull ailerons out and down.
e. Reverse the preceding steps for reinstallation.
NOTE
If rigging was correct and push pull tube
adjustment was not disturbed, it should
not be necessary to re-rig system.

SHOP NOTES:

6-8

Change 3

f. Check aileron travel and alignment, re-rig if
necessary, in accordance with paragraph 6-17.
6-16. REPAIR. Aileron repair may be accomplished
in accordance with instructions outlined in Section 18.
Before installation, ensure balance weights and hinges
are securely attached.
6-17. RIGGING.
a. (Refer to figure 6-1.) Remove access plates
and upholstery as required.
b. Remove safety wire and relieve cable tension
at turnbuckles (9 and 13).
c. (Refer to figure 6-3.) Disconnect push-pull
tubes (9) at ailerons (2).
d. (Refer to figure 6-2.) Adjust turnbuckle (37)
and adjustment nuts (35) on interconnect cable (38)
to remove slack, acquire proper tension (30±10
pounds) and position both control wheels level (synchronized).
e. Tape a bar across both control wheels to hold
them in neutral position.
f. (Refer to figure 6-1.) Adjust direct cable turnbuckles (9) and carry-thru cable turnbuckle (13) to
position bellcranks (index 19, figure 6-3) approximately in neutral while maintaining proper cable
tension.
g. Streamline ailerons with reference to flaps
(flaps full UP positions), then adjust push-pull
tubes (index 9. figure 6-3) to fit and install.
h. With ailerons streamlined, mount an inclinometer on trailing edge of aileron and set pointer to
0°.
Remove bar from control wheels and adjust
i.
travel stops (index 11, figure 6-3) to obtain travel
specified in figure 1-1.
j. Ensure all turnbuckles are safetied, all cables
and cable guards are properly installed, all jam nuts
are tight and replace all parts removed for access.

WARNING
Be sure ailerons move in correct direction
when operated by the control wheels.

SECTION 7
WING FLAP CONTROL SYSTEM

TABLE OF CONTENTS

Page

7-1
......
WING FLAP CONTROL SYSTEM
7-1
Description ..............
7-1
Operational Check ...........
7-2
Trouble Shooting ............
Flap Motor and Transmission Assembly .7-3
7-3
......
Removal and Installation
.
7-3
...........
Repair .
7-3
Flap Control Lever ..............
7-3
Removal and Installation ......
7-5
Drive Pulleys .......
7-5
.....
Removal and Installation .
Repair .............
7-5

7-1. WING FLAP CONTROL SYSTEM.
figure 7-1.)

(Refer to

7-2. DESCRIPTION. The wing flap control system
consists of an electric motor and transmission assembly, drive pulleys, synchronizing push-pull tubes,
bellcranks, push-pull rods, cables, pulleys and a
follow-up control. Power from the motor and transmission assembly is transmitted to the flaps by a
system of drive pulleys, cables and synchronizing
tubes. Electrical power to the motor is controlled by
two microswitches mounted on a "floating" arm, a
control lever and a follow-up control. As the control
lever is moved to the desired flap setting, a switch is
tripped actuating the flap motor. As the flaps move,
the floating arm is rotated by the follow-up control
until the active switch clears the control lever cam,
breaking the circuit. To reverse the direction of
flap travel, the control lever is moved in the opposite
direction. When the control lever cam contacts the
second switch the flap motor is energized in the opposite direction. Likewise, the follow-up control
moves the floating arm until the second switch is
clear of the control lever cam.

Bellcranks ..............
..
Removal and Installation
Repair ..............
Flaps.
Removaland Installation ......
Repair ..............
.......
Cables and Pulleys .....
Removal and Installation ......
.........
Rigging - Flaps
Rigging - Flap Control Lever and
.
......
.
Follow-Up .

...

7-5
7-5
7-5
7-5
7-5
7-5
7-5
7-5
7-5
7-13

motor should NOT continuously freewheel at travel
extremes.
c. BEGINNING WITH AIRCRAFT SERIAL U20601674 AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH FIGURE 7-2 SHEET 3 Check for
positive shut-off of motor at the flap travel extremes,
FLAP MOTOR MUST STOP OR DAMAGE WILL RESULT.
d. Check flaps for sluggishness in operation. In
at 110 MPH (THRU AIRCRAFT SERIALS P206flight
00648 AND U20601700) and 120 MPH (BEGINNING
WITH AIRCRAFT SERIAL U20601701). indicated
airspeed, flaps should fully extend in approximately
15.5 seconds and retract in approximately 7.5 seconds. On the ground, with engine running, the flaps
should extend in approximately 8 seconds and retract
in approximately 7.5 seconds.
e. With flaps full UP, mount an inclinometer on one
flap and set to 0 ° . Lower flaps to full DOWN position
and check flap angle as specified in figure 1- 1. Check
approximate mid-range percentage setting against
degrees as indicated on inclinometer. Repeat the
same procedure for the opposite flap.
NOTE

7-3. OPERATIONAL CHECK.
a. Operate flaps through their full range of travel,
* observing for uneven or jumpy motion, binding and
lost motion in the system. Ensure flaps are moving
together through their full range of travel.
b. THRU AIRCRAFT SERIALS P20600648 AND
U20601673 WHEN NOT MODIFIED IN ACCORDANCE
WITH FIGURE 7-2, SHEET 3. Check for positive
shut-off of motor at the flap travel extremes, the

An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
f. Remove access plates and attempt to rock drive
pulleys and bellcranks to check for bearing wear.
g. Inspect flap rollers and tracks for evidence of
binding and defective parts.

Change 1

7-1

7-4.

TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraphs 7-21 and 7-22.
TROUBLE

BOTH FLAPS FAIL TO MOVE.

BINDING IN SYSTEM AS FLAPS
ARE RAISED AND LOWERED.

LEFT FLAP FAILS TO MOVE.

FLAPS FAIL TO RETRACT.

7-2

PROBABLE CAUSE

REMEDY

Popped circuit breaker.

Reset and check continuity.
Replace breaker if defective.

Defective switch.

Place jumper across switch.
Replace switch if defective.

Defective motor.

Remove and bench test.
Replace motor if defective.

Broken or disconnected wires.

Run continuity check of wiring.
Connect or repair wiring as
necessary.

Disconnected or defective
transmission.

Connect transmission. Remove,
bench test and replace transmission if defective.

Defective limit switch.

Check continuity of switches.
Replace switches found defective.

Follow-up control disconnected or slipping.

Secure control or replace
if defective.

Cables not riding on pulleys.

Open access plates and observe
pulleys. Route cables correctly
over pulleys.

Bind in drive pulleys.

Check drive pulleys in motion.
Replace drive pulleys found
defective.

Broken or binding pulleys.

Check pulleys for free rotation or
breaks. Replace defective pulleys.

Frayed cable.

Check condition of cables.
defective cables.

Flaps binding on tracks.

Observe flap tracks and rollers.
Replace defective parts.

Disconnected or broken cable.

Check cable tension.
Connect or replace cable.

Disconnected push-pull rod.

Attach push-pull rod.

Disconnected or defective
UP limit switch.

Check continuity of switch.
Connect or replace switch.

Replace

7-4.

TROUBLE SHOOTING (Cont).
PROBABLE CAUSE

TROUBLE

REMEDY

FLAPS FAIL TO EXTEND.

Disconnected or defective
DOWN limit switch.

Check continuity of switch.
Connect or replace switch.

INCORRECT FLAP TRAVEL.

Incorrect rigging.

Refer to paragraphs 7-21 and 7-22.

Defective limit switch.

Check continuity of switches.
Replace switches found defective.

7-5. FLAP MOTOR AND TRANSMISSION ASSEMBLY.
7-6. REMOVAL AND INSTALLATION.
a. THRU AIRCRAFT SERIALS P20600648 AND
U20601673 WHEN NOT MODIFIED IN ACCORDANCE
WITH SK150-37 AND WHEN NOT MODIFIED IN ACCORDANCE WITH FIGURE 7-2, SHEET 3. (Refer
to figure 7-2, sheet 1.)
1. Run flaps to full DOWN position.
2. Disconnect battery cables at the battery and
insulate cable terminals as a safety precaution.
3. Remove access plates adjacent to drive pulley
and motor assembly on right wing.
NOTE
Remove motor (1), transmission (4), hinge
assembly (2) and actuating tube (8) from the
aircraft as a unit.
4. Remove bolt (20) securing actuating tube (8)
to drive pulley (16).
5. Screw actuating tube (8) IN toward transmission (4) by hand to its shortest length.
6. Remove bolt (3) securing flap motor hinge assembly (2) to wing, or remove bolt (5) securing transmission (4) to hinge assembly (2). Retain brass
washer between lower end of hinge and wing structure.
Remove hinge assembly (2) through access opening,
using care not to drop bushing from hinge. Tape open
ends of hinge to protect bearings.
7. Disconnect motor electrical wiring (21) at
quick-disconnects.
8. Using care, work assembly from wing through
access opening.
9. Reverse the preceding steps for reinstallation.
If the hinge (2) was removed from the transmission for
any reason, ensure the short end of hinge is reinstalled toward the top.
10. Complete an operational check as outlined in
paragraph 7-3 and re-rig flap system in accordance
with paragraphs 7-21 and 7-22.
b. THRU AIRCRAFT SERIALS P20600648 AND
U20601673 WHEN MODIFIED IN ACCORDANCE
WITH SK150-37 AND WHEN NOT MODIFIED IN
ACCORDANCE WITH FIGURE 7-2, SHEET 3.
(Refer to figure 7-2, sheet 2.)

1. Complete steps 1, 3 and 4 of subparagraph
"a."

2. Run flap motor to place actuating tube (8) IN
to its shortest length.
3. Complete steps 2, 6, 7, 8, 9 and 10 of subparagraph "a."
c. BEGINNING WITH AIRCRAFT SERIAL U20601674 AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH FIGURE 7-2, SHEET 3. (Refer to
figure 7-2, sheets 2 and 3.)
1. Complete steps 1 thru 7 of subparagraph "a."
2. Disconnect electrical wiring at limit switches
(31 and 34). Tag wires for reference on reinstallation.
3. Complete steps 8, 9 and 10 of subparagraph
"a."
7-7. REPAIR. Repair consists of replacement of
motor, transmission, coupling, actuating tube and
associated hardware. Bearings in hinge assembly
may also be replaced. Lubricate as outlined in Section 2.
7-8.

FLAP CONTROL LEVER.

7-9. REMOVAL AND INSTALLATION.
a. THRU AIRCRAFT SERIALS P20600648 AND
U20601700. (Refer to figure 7-3, sheet 1.)
1. Remove follow-up control (1) from switch
mounting arm (14).
2. Remove flap operating switches (11 and 13)
from switch mounting arm (14). DO NOT disconnect
electrical wiring at switches.
3. Remove knob (9) from control lever (8).
4. Remove remaining items by removing bolt
(17). Use care not to drop parts into tunnel area.
5. Reverse the preceding steps for reinstallation.
Do not overtighten bolt (17) causing lever (8) to bind.
Rig system in accordance with paragraphs 7-21 and
7-22.
b. BEGINNING WITH AIRCRAFT SERIAL U20601701. (Refer to figure 7-3, sheet 2 and 3.)
1. Disconnect follow-up control bellcrank (24)
from switch mounting arm (8).
2. Remove flap operating switches (15 and 16)
from switch mounting arm (8). DO NOT disconnect
electrical wiring at switches.

Change 3

7-3

NOTE

4

Shaded pulleys are used for
this system.

2

1

REFER TO FIGURE 7-2

DetailB
DetaiL

7

A

'/'

-.

.,

6

. ..r''

/·..
' ""

Detail

.. .

-'

REFER TO FIGURE 7-3

1""
· ..

.' .

4

1

/'A

-

X' . .

........

Detail D

Pulley

1.

Cable Guard
Spacer
Bushing
Bracket
Rear Carry-Thru Spar
Synchronizing Push-Pull Tube
Follow-Up Control
Rub Strip
Turnbuckle

i
19

17

"

Detail A

|CAUTIONt

Push-Pull Rod
Attach Bracket
Bearing
Support
Bolt
Washer
Bushing

MAINTAIN PROPER CONTROL
CABLE TENSION.

CABLE TENSION:
70 LBS ± 10 LBS (AT AVERAGE TEMPERATURE FOR THE AREA)
REFER TO FIGURE 1-1 FOR TRAVEL.

Figure 7-1.
Change I

'"...'.i

.,

11. Bolt
12. Bellcrank
13. Bolt

7-4

TO

-'":;.*-';;;;

\

;B.*----::
' :-: I

14.
15.
16.
17.
18.
19.
20.

'.
"REFER

...""-*'""

^

1.
2.
3.
4.
5.
6.
7.
8.
9.
10.

BB

Wing Flap Control System

3.
4.

Remove knob (11) from control lever (12).
Remove remaining items by removing bolt
(18). Use care not to drop parts into tunnel area.
5. Reverse the preceding steps for reinstallation. Do not overtighten bolt (18) causing lever (12)
to bind. Rig system in accordance with paragraphs
7-21 and 7-22.
7-10.

DRIVE PULLEYS.

(Refer to figure 7-2.)

7-11. REMOVAL AND INSTALLATION.
a. Remove access plates adjacent to drive pulley
(16) in right wing.
b. Unzip or remove headliner as necessary for
access to turnbuckles (index 10, figure 7-1), remove
safety wire and loosen turnbuckles.
c. Remove bolt (18) securing flap push-pull rod (14)
to drive pulley (16).
d. Remove bolt (10) securing synchronizing pushpull tube (9) to drive pulley (16) and lower RIGHT
flap gently.
e. Remove bolt (20) securing actuating tube (8) to
drive pulley (16) and lower LEFT flap gently. Retain
bushing.
f. Remove cable locks (13) securing control cables
to drive pulley (16). Tag cables for reference on
reinstallation.
g. THRU AIRCRAFT SERIALS P20600648 AND
U20601700. Remove bolt (11) attaching follow-up
control bellcrank (17) to drive pulley (16).
h. Remove bolt (12) attaching drive pulley (16) to
wing structure.
i. Using care, remove drive pulley through access
opening, being careful not to drop bushing. Retain
brass washer between drive pulley and wing structure
for use on reinstallation. Tape open ends of drive
pulley after removal to protect bearings.
j. To remove left wing drive pulley, use this same
procedure omitting steps "e" and "g."
k. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraphs 7-21 and
7-22, safety turnbuckles and reinstall all items removed for access.
7-12. REPAIR. Repair is limited to replacement of
bearings. Cracked, bent or excessively worn drive
pulleys must be replaced. Lubricate drive pulley
bearings as outlined in Section 2.
7-13.

BELLCRANKS.

(Refer to figure 7-1.)

7-14. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position.
b. Remove access plate adjacent to bellcrank (12).
c. Remove bolt (18) securing outboard push-pull
rod (14) to bellcrank (12).
d. Remove bolt (11) securing synchronizing pushpull tube (7) to bellcrank (12).
e. Remove bolts (13) securing upper and lower
* Supports (17).
f. Work bellcrank out through access opening.
g. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraphs 7-21 and

7-15. REPAIR. Repair is limited to replacement of
bearings. Cracked, bent or excessively worn bellcranks must be replaced. Lubricate bearings as
outlined in Section 2.
7- 16.

FLAPS.

(Refer to figure 7-4.)

7-17. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position.
b. Remove access plates (5) from top leading edge
of flap.
c. Disconnect push-pull rods at flap brackets (4).
d. Remove bolts (12) at each flap track, pull flap
aft and remove remaining bolt. As flap is removed
from wing, all washers, rollers and bushings will
fall free. Retain these for reinstallation.
e. Reverse the preceding steps for reinstallation.
If push-pull rod adjustment is not disturbed, rerigging of system should not be necessary. Check
flap travel and rig in accordance with paragraphs 7-21
and 7-22, if necessary.
7-18. REPAIR. Flap repair may be accomplished
in accordance with instructions outlined in Section
18.
7-19.
7-1.)

CABLES AND PULLEYS.

(Refer to figure

7-20. REMOVAL AND INSTALLATION.
a. Remove access plates, fairings, headliner and
upholstery as necessary for access.
b. Remove safety wire, relieve cable tension, disconnect turnbuckles (10) and carefully lower LEFT
flap.
c. Disconnect cables at drive pulleys, remove cable guards and pulleys as necessary to work cables
free of aircraft.
NOTE
To ease routing of cables, a length of wire
may be attached to the end of cable being
withdrawn from the aircraft. Leave wire
in place, routed through structure; then attach the cable being installed and use wire
to pull cable into position.
d. Reverse the preceding steps for reinstallation.
e. After cables are routed in position, install pulleys and cable guards. Ensure cables are positioned
in pulley grooves before installing guards.
f. Re-rig flap system in accordance with paragraphs
7-21 and 7-22, safety turnbuckles and reinstall all
items removed in step "a."
7-21. RIGGING-FLAPS. (Refer to figure 7-2.)
a. Unzip or remove headliner as necessary for access to turnbuckles (index 10, figure 7-1).
b. Remove safety wire, relieve cable tension, disconnect turnbuckles and carefully lower LEFT flap.
c. Remove bolt (18) securing flap push-pull rod
(14) to drive pulleys (16) in both wings.

7-22.

Change 1

7-5

1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11. Bolt
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.

Motor Assembly
Hinge Assembly
Pivot Bolt
Transmission Assembly
Bolt
Nut and Ball Assembly
Setscrew
Actuating Tube
Synchronizing Push-Pull Tube
Bolt
Pivot Bolt
Cable Lock
Push-Pull Rod
Attach Bracket
Drive Pulley
Bellcrank
Bolt
Spacer
Bolt
Electrical Wiring
Down-Limit Switch
Up-Limit Switch
Snubber Assembly
Bracket
Spacer
Shim
Screw
Setscrew
Switch Adjusting Block
Up-Limit Switch
Switch Actuating Collar
Switch Support
Down-Limit Switch
Bushing
Figure 7-2.

7-6

Change 1

Use Loctite Sealant, Grade "C" on
threads of setscrew (7) after final
adjustment.
Ensure shortest end of hinge (2) is at top.
Flap Motor and Transmission Assembly (Sheet 1 of 3)

THRU AIRCRAFT SERIALS p20600648 AND U20601673
WHEN MODIFIED IN ACCORDANCE WITH SK150-37

4

A

31

fl

UP

7

VIEW

aps in the full
position.

A-A

BEGINNING WITH AIRCRAFT

7-7
7 -7

4

A

8

34

33

32

.12± .05 " with flaps
in the full UP position.

31

VIEW A-A

THIS FLAP ACTUATOR INSTALLATION IS EFFECTIVE
THRU AIRCRAFT SER IALS P20600648 AND U20601673
WHEN USED AS A REPLACEMENT SPARE FOR SK15037 OR PRODUCTION FLAP ACTUATOR INSTALLATIONS
PRIOR TO U20601674

Figure 7-2.
7-8

Flap Motor and Transmission Assembly (Sheet 3 of 3)

d. Remove bolt (10) securing synchronizing pushpull tube (9) to drive pulley (16) in right wing and
carefully lower RIGHT flap.
e. Remove bolt securing synchronizing push-pull
tube to drive pulley in left wing.
f. Disconnect outboard flap push-pull rods from
bellcranks in both wings.
g. Disconnect actuating tube (8) from drive pulley
(16).
NOTE
Ensure that the 3/32 inch retract cable is
connected to the forward side of the right
drive pulley and to the aft side of the left
drive pulley and that the 1/8 inch direct
cable is connected to the aft side of the
right drive pulley and to the forward side
of the left drive pulley. Ensure that the
right drive pulley rotates clockwise, when
viewed from below, as the flaps are ex-and
tended. ( Refer to figure 7-5. )NOT

align RIGHT drive pulley so that the centerline of bolt
hole for inboard push-pull rod is 4.20 inches aft of
fuel well bulkhead (refer to figure 7-5). Tighten setscrew (7) in accordance with procedures outlined in
the following note and secure actuating tube to drive
pulley with bolt (20).
NOTE
Thru Aircraft Serial U20602223: Tighten
setscrew (7). Aircraft Serials U20602224
thru U20602376: Apply grade CV sealant
to setscrew (7) threads and torque to 45
lb-in. Beginning with Aircraft Serial
U20602377: Apply grade CV sealant to
setscrew (7) threads and torque to 60 lb-in.
1. Manually holding RIGHT flap full up, adjust
push-pull rods to align with drive pulley and bellcrank attachment holes. Connect push-pull rods
locknuts.

h. Adjust synchronizing push-pull tube (9) in RIGHT
The right flap and actuator MUST be correctly
MUST be correctly
actuator
andand
right flap
The before
left flap can be rigrigged
cables
wing to 48.69 inches between centers of rod end holes,
rigged before cables and left flap can be rigged
tighten jam nuts and connect to bellcrank and drive
pulley.
i. THRU AIRCRAFT SERIALS P20600648 AND
m. Mount an inclinometer on trailing edge of RIGHT
U20601673 WHEN NOT MODIFIED IN ACCORDANCE
WITH SK150-37 AND WHEN NOT MODIFIED IN
ACCORDANCE WITH FIGURE 7-2, SHEET 3. (ReferNOTE
to figure 7-2, sheet 1.) Screw actuating tube (8) IN
toward transmission (4) by hand to its shortest length
An inclinometer for measuring control surface
(flaps full up position). Loosen setscrew (7) securing
trael is available from the Cessna Service
actuating tube to nut and ball assembly (6), hold nut
Parts Center. Refer to figure 6-4.
and ball assembly so that it will not move and adjust
actuating tube IN or OUT as necessary to position the
n. THRU AIRCRAFT SERIALS P20600648 AND
RIGHT drive pulley so that the centerline of bolt hole
U20601673 AND ALL AIRCRAFT NOT MODIFIED
for the inboard push-pull rod attachment is 4.20
IN ACCORDANCE WITH FIGURE 7-2, SHEET 3.
inches aft of fuel well bulkhead (refer to figure 7-5).
1. With RIGHT flap in full UP position, adjust
Tighten setscrew (7) and secure actuating tube to
UP-LMIT switch (23) to operate and shut-off elecdrive pulley with bolt (20).
trical power to motor at degree of travel specified
j. THRU AIRCRAFT SERIALS P20600648 AND
U20601673 WHEN MODIFIED IN ACCORDANCE WITH
1-1.
in figure
2.
flap to to
DOWN position
and
and adjust
Run RIGHT
RIGHT
ACCOR2. Run
AND WHEN
SK150-37
SK150-37 AND
WHEN NOT
NOT MODIFIED
MODIFIED IN
IN ACCORDOWN-LIMIT
switchflap
(22) toDOWN
operateposition
and shut-off
DANCE WITH
WITH FIGURE
FIGURE 7-2,
7-2, SHEET
SHEET 3.
3. Operate
Operate flap
flap
DOWN-LIMIT
switch
(22) to
and
shut-off
DANCE
electrical
power
to motor
at operate
degree of
travel
specielectrical
power to motor at degree of travel specmotor until actuating tube (8) is IN to its shortest
length (flaps full up position). Loosen setscrew (7)
fied
in figure 1-1.
o. BEGINNING WITH AIRCRAFT SERIAL U206U206- BEGINNNG WITH AIRCRAFT Serial
securing actuating tube to nut and ball assembly (6),
01674 AND ALL AIRCRAFT MODIFIED IN ACCORhold nut and ball assembly so that it will not move
DANCE WITH FIGURE 7-2, SHEET 3.
and adjust actuating tube IN or OUT as necessary to
position the RIGHT drive pulley so that the centerline
1. With RIGHT flap in full UP position, loosen
setscrew (29) and slide UP-LIMIT switch (31) adjustof bolt hole for the inboard push-pull rod attachment
ment block (30) to operate switch and shut-off elecis 4.20 inches aft of fuel well bulkhead (refer to figure
trical power to motor at degree of travel specfied in
7-5). Tighten setscrew (7) and secure actuating tube
to drive pulley with bolt (20).
figure 1-1. Tighten setscrew (29).
2. Run RIGHT flap to DOWN position and adjust
k. BEGINNING WITH AIRCRAFT SERIAL U206DOWN-LIMIT switch (34) adjustment block (30) to
01674 AND ALL AIRCRAFT MODIFIED IN ACCORoperate switch and shut-off electrical power to motor
at degree of travel specified in figure 1-1. Tighten
ating tube (8) IN toward transmission (4) by hand to
setscrew (29).
12±.05 inches between switch actuating collar (32)
p. Run RIGHT flap to full UP position.
and transmission as illustrated in figure 7-2, VIEW q.
Complete step "h" for synchronizing push-pull
A-A. Loosen setscrew (7) securing actuating collar
tube in LEFT wing
(32). Hold actuating collar to maintain . 12±.05" and
r. Connect control cables at turnbuckles (index 10,
adjust actuating tube (8) IN or OUT as necessary to
figure 7-1). Adjust turnbuckles to position left drive
pulley so that the centerline of bolt hole for the inChange 2

7-9

1. Follow-Up Control
2. Bracket
3. Clamp
4. Spacer
5. Washer

6. Spring
7. Cam
8. Control Lever

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

Knob
Bracket
Flaps DOWN Operating Switch
Insulator
Flaps UP Operating Switch
Switch Mounting Arm
Flap Position Indicator
Bushing
Bolt
Stiffener

'I

of control

...

lever

(8)
Istalletion
upon

A

*.18
$>

i

:

-

d

/

/ .

17
'"'"/ '

5'":*4"":®"

"
;

'i"/=i >~lJ 1~~ ·

11
1/ S/ 13 / '^'LV
*10ol /

It
**12
Detail A

Ax\

'<,\[AND

Figure 7-3.

7-10

Change 3

1

/ .-. />' /:

r

.*
|

~Apply

~-^

jiJ

*

NOTE

Grade "C" Loctite to threads
of control lever (8) upon installation
of knob (9).

BEGINNING WITH AIRCRAFT SERIALS P206-0534
AND U206-1237 THRU U20601590

BEGINNING WITH AIRCRAFT SERIAL U20601633

as_.<

** BEGINNING WITH AIRCRAFT SERIALS P206-0557
U206-1248

Flap Control Lever Installation (Sheet 1 of 3)

__

~~~~~~A

-~

/~1

28 0

2

290

EyX*4

t

N/

..
A^-$

/ 1 ?S

U20601923

/10

Detail A
e

1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.

BEGINNING WITH SERIALS
& ON.

.1
DU20601924

B

Follow-Up Control
Control Adjustment Nut
Control Bracket
Nylon Guide
Union Assemblyt
Retract Cable
Bushing
Arm Assembly
Flap Position Indicator
Bracket
Knob1
Control Lever
Washer
Cam
Flaps UP Operating Switch
Flaps DOWN Operating Switch

18. Bolt
19. Spacer
20. Bolt
21. Spring
22. Firewall Stiffener
23. Bracket
24. Bellcrank Assembly
25. Pin
26. Bracket
27. Pedestal Structure
28. Metal Washer
29. Nylon Washer
30. Support

6

THRU SERIAL

""~/ , ^,.
'< *,/--

/"

/

A

Detail

-29?

'

29290S

*

8

//

/

1i5
<7

.16
/
20
22

1

21

;

>\p
26

1

25
\

\

/

27

13

/27/

/

t

21
nt
NOTE

/

Apply Grade "C" Loctite to threads
of control lever (12) upon installation of knob (11).

* THRU AIRCRAFT SERIALS P20600648 AND U20601912
* BEGINNING WITH AIRCRAFT SERIAL U20601913
( Refer to Cessna Service Letter SE 73-8 for additional information ).

Figure 7-3.

B

Detail
SERIALS U20601701 THRU
U20603020

_

Flap Control Lever Installation (Sheet 2 of 3)
Change 3

7-11

1. Follow-Up Control

2. Control Adjustment Nut
3. Control Bracket
4. Nylon Guide
5. Union Assembly
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25. Pin
26.
27.
28.
29.
30.

Retract Cable
Bushing
Arm Assembly
Flap Position Indicator
Bracket
Knob
Control Lever
Washer
Cam
Flaps UP Operating Switch
Flaps DOWN Operating Switch
Insulator
Bolt
Spacer
Bolt
Spring
Firewall Stiffener
Bracket
Bellcrank Assembly
Bracket
Pedestal Structure
Metal Washer
Nylon Washer
Support

Change 3

7

BEGINNING WITH SERIAL
U20603021

Figure 7-3.
7-12

20

Flap Control Lever Installation (Sheet 3 of 3)

A
B

C

Detail

A

BEGINNING WITH AIRCRAFT

8 9 13

SERIALS P206-0520 AND U206-

/ /

1235 THRU U20601568

BEGINNING WITH AIRCRAFT
SERIAL U20601569/
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.

Wing Structure
Flap Track
Wing Flap
Bracket
Access Plate
Nylon Plug Button
Stiffener
Nut
Washer
Roller
Bushing
Bolt
Spacer

10

1

/
Detail C
12
Detail

B
NOTE
Bushings (11), rollers (10) and spacers (13) are first
positioned through slots in flap tracks, then are secured to the flap roller supports with attaching bolts,
washers and nuts. Nylon plug buttons (6) prevent flap
from chafing wing trailing edge.

Figure 7-4.

Flap Installation
Change 3

7-12A/(7-12B blank)

DRIVE PULLEY

FWD

SYNCHRONIZING
PUSH-PULLTUBE

FLAPMOTOR AND
TRANSMISSION

FUEL WELL
BULKHEAD

SET SCREW

4 . 24.20
0

OUTBOARD
PUSH-PULL
ROD

INBOARD
PUSH-PULL
PUSH-PULL
ROD

LEFT WING

FLAP MOTOR
ACTUATING
TUBE
TUBE
BELLCRANK

TURNBUCKLES

VIEWED FROM ABOVE

Figure 7-5.

RIGHT WING

Flap System Schematic

board push-pull rod attachment is 4.20 inches aft of
fuel well bulkhead, maintaining 70±10 pounds tension.
Adjust retract cable first.
NOTE
Ensure cables are positioned in pulley grooves
and cable ends are positioned correctly at
drive pulleys before tightening turnbuckles.

maintaining this position.
3. Mount an inclinometer on trailing edge of
one flap and set to 0° . Turn master switch ON and
move control lever to the 10° position. If flap travel
is more than 10 ° , adjust flaps DOWN operating switch
(11) away from cam (7) and recycle flaps. If flap
travel is less than 10 ° , adjust flaps DOWN operating
switch (11) closer to cam (7) and recycle flaps.
NOTE

s. Manually holding LEFT flap full UP, adjust
push-pull rods to align with drive pulley and bellcrank attachment holes. Connect push-pull rods
and tighten locknuts.
t. After completion of steps "a" thru "s", operate
flaps and check for positive shut-off of flap motor
through several cycles. Check for specified flap
travel with inclinometer mounted on each flap separately.
NOTE
Since the flap rollers may not bottom in the
flap tracks with flaps fully extended, some
free play may be noticed in this position.
7-22. RIGGING-FLAP CONTROL LEVER AND
FOLLOW-UP.
a. THRU AIRCRAFT SERIALS P20600648 AND
U20601700. (Refer to figure 7-3, sheet 1.)
1. Disconnect follow-up control rod end (1) at
switch mounting arm (14).
2. Move control lever (8) to full UP position,
then without moving control lever, move switch
mounting arm (14) until cam (7) is centered between
switches (11 and 13). Adjust follow-up control rod
end (1) to align with the attaching hole in the switch
mounting arm and secure rod end to mounting arm

An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
4. Adjust flaps UP operating switch (13) in slotted holes for .062 inch clearance between switch roller and cam (7) when the flaps DOWN operating switch
has just opened in the 10° and 20° position.
NOTE
Flap travel on UP cycle may deviate a maximum of 4° from indicated position.
5. Turn master switch ON and run flaps through
several cycles, stopping at various mid-range settings and checking that cable tension is within limits.
Retract cable tension may increase to 90 pounds when
flaps are fully retracted.
6. Check all rod ends and clevis ends for sufficient thread engagement, all jam nuts are tight and
reinstall all items removed for access.
7. Flight test aircraft and check that follow-up
control does not cause automatic cycling of flaps. If
cycling occurs, readjust operating switches as necessary per steps 2, 3 and 4.
Change 2

7-13

b. BEGINNING WITH AIRCRAFT SERIAL U20601701.
(Refer to figure 7-3, sheet 2 and 3. )
1. Run flaps to full UP position.
2. Remove upholstery and headliner as necessary for access.
3. Pull all slack from follow-up control cable
and with position indicator (9) in the full UP position,
secure follow-up cable to retract cable (6) with union
assembly (5). Ensure union assembly is at end of
support (30).
4. Connect spring (21) to bellcrank (24).
5. Make minor cable length adjustments at
brackets (3) by adjusting nuts (2).
6. With control lever (12) in lull UP position,
adjust switches (15 and 16) in slotted holes until cam
(14) is centered between switch rollers. Be sure
control lever (12) is in full UP position during this
adjustment.
7. Mount an inclinometer on trailing edge of one
flap and set to 0 ° . Turn master switch ON and move
control lever to 10 ° position. If flap travel is more
than 10 ° , adjust flaps DOWN operating switch (16)
away from cam (14) and recycle flaps. If flap travel
is less than 10 ° , adjust flaps DOWN operating switch

SHOP NOTES:

7-14

Change 3

(16)
closer to cam (14) and recycle flaps.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
8. Adjust flaps UP operating switch (15) in slotted holes for .062 inch clearance between switch roller and cam (14) when the flaps DOWN operating
switch has just opened in the 10° and 20 ° position.
NOTE
Flap travel on UP cycle may deviate a maximum of 4° from indicated position.
9. Turn master switch ON and run flaps through
several cycles, stopping at various mid-range settings and checking that cable tension is within limits.
Retract cable tension may increase to 90 pounds when
flaps are fully retracted.

SECTION 8
ELEVATOR CONTROL SYSTEM

Page

TABLE OF CONTENTS
ELEVATOR CONTROL SYSTEM .......
Description (Thru U20602579) ......
Description (Beginning with U20602580). .
..
Trouble Shooting ... . . . .. .
.
Control Column ...........
.
. . . . . . ...
......
Elevators .
........
Removal and Installation ..........
Repair ..............

8-1. ELEVATOR CONTROL SYSTEM.
U20602579) (Refer to figure 8-1.)

. 8-7
.........
Bellcrank ....
8-7
Removal and Installation ......
8-7
Arm Assembly .......
8-7
Removal and Installation ......
....
8-7
.
...
Cables and Pulleys
8-7
Removal and Installation ......
. 8-8
.............................
Rigging (Thru U20602579).
8-9
Rigging (Beginning with U20602580) ...

8-1
8-1
8-1
8-1
8-2
8-2
8-2
8-7

8-2A. ELEVATOR CONTROL SYSTEM BEGINNING
WITH AIRCRAFT SERIAL U2062580. (Refer to
figure 8-1 A.)

(THRU

8-2. DESCRIPTION. The elevators are operated by
power transmitted through fore-and-aft movement of
the pilot or copilot control wheels. The system is
comprised of control columns, an elevator torque
tube, cables and pulleys. The elevator control cables,
at their aft ends, are attached to a bellcrank mounted
on a bulkhead in the tailcone. A push-pull tube connects this bellcrank to the elevator arm assembly, installed between the elevators. An elevator trim tab
is installed in the trailing edge of the right elevator
and is described in Section 9.

8-3.

8-2B. DESCRIPTION. Beginning with aircraft serial
U20602580 and on. the single large elevator down
spring is replaced by two smaller springs which
attach to each side of the elevator bellcrank and anchor to the lower forward face of the tailcone bulkhead. The elevator up and down cables are re-routed
from the elevator control arm assembly through the
fuselage to the elevator bellcrank in the tailcone. The
elevator up cable is routed to the top turnbuckle connected to the elevator bellcrank.

TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 8-14.

TROUBLE
NO RESPONSE TO CONTROL
WHEEL FORE-AND-AFT
MOVEMENT.

PROBABLE CAUSE

REMEDY
Attach push-pull

Forward or aft end of push-pull
tube disconnected.

Check visually.
tube correctly.

Cables disconnected.

Check visually. Attach cables and
rig system in accordance with
paragraph 8- 14.

Change 3

8-1

8-3.

TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

BINDING OR JUMPY MOTION
FELT IN MOVEMENT OF ELEVATOR SYSTEM.

ELEVATORS FAIL TO ATTAIN
PRESCRIBED TRAVEL.

Defective bellcrank or arm
assembly pivot bearings or
push-pull tube attach bearings.

Move bellcrank or arm to check for
play or binding. Disconnect pushpull tube and check that bearings
rotate freely. Replace defective
parts.

Cables slack.

Check and adjust to tension specified
in figure 8- 1.

Cables not riding correctly on
pulleys.

Check visually. Route cables correctly over pulleys.

Nylon grommet on instrument
panel binding.

Replace grommet.

Defective control column
bearing rollers.

Check visually.
rollers.

Defective control column
torque tube bearings.

Disconnect necessary items and
check that bearings rotate freely.
Replace defective bearings.

Control guide on aft end of
control square tube adjusted
too tightly.

Loosen screw and tapered plug
in end of control tube enough to
eliminate binding.

Defective elevator hinges.

Disconnect push-pull tube and
move elevators by hand. Replace
defective hinges.

Defective pulleys or cable
guards.

Check visually. Replace defective
parts and install guards properly.

Stops incorrectly set.

Rig in accordance with paragraph 8-14.

Cables tightened unevenly.

Rig in accordance with paragraph 8-14.

Interference at instrument
panel.

Rig in accordance with paragraph 8-14.

8-4. CONTROL COLUMN. (Refer to figure 6-2.)
Section 6 outlines removal, installation and repair of
control column.
8-5.

ELEVATORS.

(Refer to figure 8-2.)

8-6. REMOVAL AND INSTALLATION.
a. Remove stinger.
b. Disconnect trim tab push-pull tube at tab actuator. (Refer to Section 9.)

8-2

Change 2

REMEDY

Replace defective

NOTE
If trim system is not moved and actuator screw
is not turned, re-rigging of trim system should
not be necessary after reinstallation of elevator.
c. Remove bolts (13) securing elevator torque tubes
(7) to arm assembly (8).
d. Remove bolts (6) from elevator hinges (5).
e. Using care, remove elevator.

REFER TO FIGURE 8-2

REFER TO
FIGURE 8-3

Detail

A

Detail

Detail C

B

* ..... .

9
ELEVATOR

DOWN

DetailD

-

B

Detail

Shaded
for this pulleys
system.are used

ELEVATOR

Detail

F

Detail

G

THRU AIRCRAFT SERIAL
P20600648 AND U20601700

/

~~~~/
BEGINNING WITH AIRCRAFT

2

*
Detail I

H

2.
3.
4.
5.
6.
7.
8
9.
10.
11.

Cable Guard
Arm Assembly
Elevator Torque Tube
Downspring
Spacer
Clip
Fairlead
Elevator Up Cable
Turnbuckle
Elevator Down Cable

SERIAL U20601701

ELEVATOR UP

ELEVATORI

MAINTAIN PROPER CONTROL

/

DOWN

CABLE TENSION.

SERIAL U20602579
Figure 8-1.

-

CABLE TENSION:
30 LBS ± 10 LBS (AT THE AVERAGE
TEMPERATURE FOR THE AREA.).
REFER TO FIGURE 1-1 FOR TRAVEL

Elevator Control System
Change 2

8-3

12

4

19

' D Lf
//--='
wDetail

7is

13

;

5\
S_^

10
g/ gg)

4.\
/_ g)

Arm Assembly
Elevator Torque Tube
Elevator Up Cable
Elevator Down Cable
Pulley
Cable Guard
Bolt
Nut
4
Cotter Pin
Turnbuckle
Elevator Down Spring
Elevator Bellcrank
Elevator Cable Link
Bearing
Push-Pull Tube
Fairlead
Clip
Washer
2_
Spacer

3

\

I

1

G
,

B

19

A

ELEVATOR

L.,
/ ;,
19

/

· 5 -,

,

1/

/
(

<

,

Figure 8-1A.
Change 2

t6

,'I
1

ELEVATOR;
DOWN

Detail

BEGINNING WITH AIRCRAFT
SERIAL U20602580

8-4

Detail G

ELEVATOR UP 5)DOWN
o

j9
Detail

FIGURE 8-3A)

A

i

3-

__REFER TO

[

D

ELEVATOR UP I

11

F

Detail
-

/

TOI
FIGURE 8-2

DetailB
B

1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.

\

.
,y

'REFER
t/

!:

2

1

9

7

Detail

C

Elevator Control System

E

A

7

13

C

Detail C

THRU AIRCRAFT SERIALS
P20600648 AND U20601874
Detail

1. Trim Tab
2. Right Elevator
3. Left Elevator
4. Balance Weight
5. Hinge Assembly
6. Bolt
7. Torque Tube
8. Arm Assembly
9. Needle Bearing
10. Bolt
11. Push-PullTube
12. Pivot Bolt
13. Bolt

BEGINNING WITH AIRCRAFT SERIAL U20601875

B

.18

3.00
THIS VIEW APPLIES TO THE RIGHT HAND ELEVATOR
WHEN THE LEFT HAND ELEVATOR IS STREAMLINED

NOTE
Do not attempt to align the elevator trailing edges
as there is a 0 ° 54' twist designed into the connecting torque tube. This twist causes the right
elevator to be higher than the left.

Figure 8-2.

Elevator Installation
Change 2

8-5

With elevators in the full down position, adjust
turnbuckle (4) and downspring (34 for an overall length of downspring to be 7.80 inches
(71±3 Ibs); safety wire turnbuckle.

9

3

Down-Spring
Turnbuckle
5. Pulley
3.
4.

6.

8

I

Down-Spring Cable
Stop Block
Bracket
Pivot Bolt
Push-Pull Tube

12.

Bolt

'i

ire DOWN
urnbuckle.
safety
loop)
ang 14.ir
C,1
18.,E"'v',oI

-

19

<
12

1
2

1I

---

THRU AIRCRAFT
SERIAL U206025'79

18

8-3.

12

7

'

Figure 8-3.
Change 2

-

A,\
17

With elevators in the full down position, adjust
turnbuckle (4) and downspring (3) to a length of
10.45 inches at 361 lbs spring tension(measured
from centerline of turnbuckle fork to outside of

8-6

,2-

;l

Cable Guard

7.
8.
9.
10.
11.

13. Bearing
14. Bellcrank Assembly
15. Link Assembly
16. Link
4. Turnbuckle
17.
Turnbuckle
18. Elevator DOWN Cable
19. Elevator UP Cable

spr-

l

0/

F
,

7

A
3

/

BEGINNING WITH
AIRCRAFT SERIAL
U20602580

-,:r
vElevmblr

Installatdon

Elevator Bellcran

Installation

\,

S'

2.

TO
*-ELEVATOR
UP CABLE

BELLCRANK
/

NOTE
BELLCRANK
STOPS

Holes are drilled off center in bellcrank
stops to provide elevator travel adjustments. 90 ° rotation of bellcrank stop
provides approximately 1° of elevator
travel.

j....
; .-;

-

Figure 8-4.

-^..t
"' \

^

ELEVATOR
PUSH-PULL
TUBE

Elevator Bellcrank Travel Stop Adjustment

8-7. REPAIR. Repair may be accomplished as outlined in Section 18. Hinge bearings may be replaced
as necessary. If repair has affected static balance,
check and rebalance as required.
BELLCRANK.

s i

TO
ELEVATOR
DOWN CABLE

f. To remove left elevator use same procedure,
omitting step "b."
g. Reverse the preceding steps for reinstallation.

8-8.

=.
Is\

(Refer to figure 8-3.)

8-9. REMOVAL AND INSTALLATION.
a. Remove access plate below bellcrank on tailcone.
(CAUTIONl|
Position a support stand under tail tie-down
ring to prevent the tailcone from dropping
while working inside.
b. Remove safety wire, relieve cable tension at
turnbuckles (17) and disconnect turnbuckle eyes at
bellcrank links (16).
c. Remove safety wire, relieve cable tension at
turnbuckle (4) and disconnect cable (7) at link assembly (15).
d. Remove bolt (12) securing push-pull tube (11) to
bellcrank (14).
e. Remove pivot bolt (10) attaching bellcrank (14)
to brackets (9) and remove bellcrank.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance wtih paragraph 8-14,
safety turnbuckles and reinstall all items removed
for access.

8-10.

ARM ASSEMBLY.

(Refer to figure 8-2.)

8-11. REMOVAL AND INSTALLATION.
a. Remove stinger.
b. Remove bolt (10) securing push-pull tube (11) to
arm assembly (8).
c. Remove bolts (13) attaching elevator torque tubes
(7) to arm assembly (8).
d. Remove pivot bolt (12) securing arm assembly
(8) and slide assembly from between elevator torque
tubes.
e. Reverse the preceding steps for reinstallation
and reinstall all items removed for access.
(Refer to figure

8-12.
8-1.)

CABLES AND PULLEYS.

8-13.

REMOVAL AND INSTALLATION.

}CAUTIONl
Position a support stand under tail tie-down
ring to prevent the tailcone from dropping
while working inside.
a. Remove seats, upholstery and access plates as
necessary.
b. Remove safety wire and relieve cable tension at
turnbuckles (10).
c. Disconnect cables at control column arm assemblies (3).
d. Disconnect cables at bellcrank links (index 16,
figure 8-3).
Change 2

8-7

1
BEGINNING WITH AIRCRAFT SERIAL U20601701

4

=

Neutral Position Dimension A

THRU AIRCRAFT SERIALS
P20600648 AND U20601700

1.30 Inches

1. Instrument Panel
2. Control Lock Collar
3. Control Lock Holes
4. Control Wheel

Figure 8-5. Control Column Neutral Rigging Position.
e. Remove fairleads, cable guards and pulleys as
necessary to work cables free of aircraft.
NOTE
To ease routing of cables, a length of wire
may be attached to the end of cable being
withdrawn from aircraft. Leave wire in
place, routed through structure; then attach the cable being installed and pull cable
into position.
f. Reverse the preceding steps for reinstallation.
g. After cables are routed in position, install pulleys and cable guards. Ensure cables are positioned
in pulley grooves before installing guards.
h. Re-rig system in accordance with paragraph
8-14, safety turnbuckles and reinstall all items removed in step "a."
8-14. RIGGING.
8-3. )

(Thru U20602579) (Refer to figure

CAUTION
Position a support stand under tail tie-down
ring to prevent tailcone from dropping while
working inside.
a. Lock control column in neutral position. (Refer
to figure 8-5)
b. Adjust turnbuckles (17) equally to streamline
LEFT elevator with horizontal stabilizer and to
obtain 30±10 lbs cable tension. (RIGHT elevator will
be higher than the left elevator) as illustrated in
figure 8-2.) Safety turnbuckles.

8-8

Change 3

NOTE
Disregard counterweight areas of elevators
when streamlining. These areas are contoured to be streamlined at cruising speed
(elevators approximately 3 ° down).
c. With elevators in the full down position, adjust
turnbuckle (4) and downspring (3) for an overall
length of downspring to be 7.80 inches (71±3 lbs);
safety wire turnbuckle (4).
d. With LEFT elevator streamlined, mount an inclinometer on elevator and set to 0°.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center (refer to figure 6-4.)
e. Adjust bellcrank travel stop blocks (8) to obtain
degree of elevator travel as specified in figure 1-1.

NOTE
Bellcrank stop blocks (8) are four-sided
bushings, drilled off-center so they may
be rotated to any one of four positions to
attain correct elevator travel. Each 90degree rotation of the stop changes the
elevator travel approximately one degree.
f. Move control wheel through full range of travel
and check cable tension in various positions. Tension should not be less than 20 pounds or more than
40 pounds in any position.

D2007C3-13 Temporary Change 2
22 February 1978

g. Check to see that all turnbuckles are safetied
and all parts are secured, then reinstall all parts
removed for access.

elevator bellcrank and elevator control cables.
d. With left elevator in streamlined position,
mount an inclinometer on elevator and set to 0°.

WARNING

NONE

Be sure elevators move in the correct direction when operated by the control wheels.

An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.

8-14A.

RIGGING.

(Beginning with U20602580).

CAUTION
Position a support stand under tail tie-down
ring to prevent the tailcone from dropping
while working inside.
a. Place contour block on left hand elevator and
lock control column in neutral position. (Refer to figure 8-5.)
b. With elevators in the full down position, adjust
turnbuckles (4) and downspring (3) to a length of
10.45 inches at 36±1 lbs spring tension (measured
from centerline of turnbuckle fork to outside of
spring loop; safety wire turnbuckles (4).
c. Install turnbuckles (4) and downsprings (3) to

e. Adjust bellcrank travel stop blocks (16) to obtain
of elevator travel as specified in figure 1-1. )
f. Move control wheel through full range of travel
and check cable tension in various positions. Tension
should not be less than 20 pounds or more than 40
pounds in any position.
g. Ensure that all turnbuckles are safetied and all
parts secured, then re-install all parts removed for
access.
WARNING
Be

sure elevators move in the correct direction when operated by the control wheels.

SHOP NOTES:

D2007C3-13 Temporary Change 2
22 February 1978

Change 3

8-9/(8-10 blank)

SECTION 9
ELEVATOR TRIM TAB CONTROL SYSTEM

|

TABLE OF CONTENTS

Page

ELEVATOR TRIM TAB CONTROL SYSTEM .
...... .
....
..
Description ...
Trouble Shooting ....
..
Trim Tab .
...........
.
Removal and Installation ......
.........
Trim Tab Actuator .
.
Removal and Installation .....
........
..
Disassembly
..
Cleaning, Inspection and Repair ..
Reassembly ......
Trim Tab Free-Play Inspection .....
Trim Tab Control Wheel ......
.
Removal and Installation......

9-1
9-1
9-1
9-2
9-2
9-2
9-2
9-2A
9-2A
9-2A
9-2A
9-4
9-4

.
9-4
Cables and Pulleys ..........
. ..
9-4
Removal and Installation ..
Pedestal Cover
............
9-7
Removal and Installation ......
9-7
. 9-7
Rigging . ...
. .............
..
. 9-8
Electric Trim Assist Installation
9-8
Description ..
. . . .......
9-8
..
Trouble Shooting . . . . ..
...
9-8
Removal and Installation ...
Clutch Adjustment ...
......
9-13
Dual Voltage Regulator Adjustment . . 9-14
Rigging ....
...
...
. 9-15

9-1. ELEVATOR TRIM TAB CONTROL SYSTEM.
(Refer to figure 9-1.)
9-2. DESCRIPTION. The elevator trim tab, located
on the trailing edge of the right elevator, is controlled by a trim wheel mounted in the pedestal. Power
to operate the tab is transmitted from the trim control wheel by means of roller chains, cables, an actuator and a push-pull tube. A mechanical pointer, ad9-3.

jacent to the trim wheel indicates nose attitude of the
aircraft. Forward rotation of the wheel trims the
nose down and aft rotation of the wheel trims the nose
up. An electric trim assist may be installed and is
described in paragraph 9-16. When de-energized
the electric trim assist has no effect on manual operation.

TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 9-14.

TROUBLE
TRIM CONTROL WHEEL MOVES
WITH EXCESSIVE RESISTANCE.

PROBABLE CAUSE

REMEDY

Cable tension too high.

Check cable tension and adjust.

Pulleys binding or rubbing.

Check pulleys visually. Repair
or replace as necessary.

Cables not in place on pulleys.

Check visually.
correctly.

Trim tab hinge binding.

Disconnect actuator and move tab
up and down to check hinge resistance. Lubricate or replace hinge
as necessary.

Defective trim tab actuator.

Remove chain from actuator
sprocket and operate actuator
manually. Replace defective
actuator.

Rusty chain.

Check visually.
rusty chain.

Install cables

Replace

Change 1

9-1

9-3.

TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

TRIM CONTROL WHEEL MOVES
WITH EXCESSIVE RESISTANCE
(CONT).

REMEDY

Damaged sprocket.

Check visually.
sprockets.

Bent sprocket shaft.

Observe motion of sprockets.
Replace defective shafts.

Cable tension too low.

Check cable tension and adjust.

Broken pulley.

Check visually.
pulley.

Replace defective

Cables not in place on pulleys.

Check visually.
correctly.

Install cables

Worn trim tab actuator.

Disconnect trim tab and check
for play in actuator. Replace
defective actuator.

Actuator attachment loose.

Check actuator for security and
tighten.

TRIM INDICATION INCORRECT.

Indicator incorrectly engaged
on wheel track.

Check visually.

INCORRECT TRIM TAB
TRAVEL.

Stop blocks loose or incorrectly
adjusted.

Adjust stop blocks on cables.
Refer to figure 9-4.

Incorrect rigging.

Refer to paragraph 9-14.

LOST MOTION BETWEEN
CONTROL WHEEL AND
TRIM TAB.

9-4.

TRIM TAB.

(Refer to figure 9-2.)

9-5. REMOVAL AND INSTALLATION.
a. Disconnect push-pull tube (9) from horn assembly (6).

NOTE
If trim system is not moved and actuator
screw is not turned, re-rigging of system
should not be necessary after reinstallation
of tab.
b. Remove screw (11) securing hinge pin (10), pull
pin until free of tab and remove tab.
NOTE
It is not necessary to completely remove
hinge pin.
c. Reverse the preceding steps for reinstallation.
Rig system,if necessary, in accordance with paragraph
9-14.

9-2

Change 1

9-6.

TRIM TAB ACTUATOR.

Replace damaged

Reset indicator.

(Refer to figure 9-1.)

9-7. REMOVAL AND INSTALLATION.
a. Relieve cable tension at turnbuckle (11).

CAUTION
Position a support stand under tail tie-down
ring to prevent tailcone from dropping while
working inside.
b. Disconnect push-pull tube (15) at actuator (19).
c. Remove access plate beneath actuator.
d. Remove chain guard (21) and disengage roller
chain (23) from actuator sprocket (20).
e. Remove screws attaching clamps (22) to bracket
(18) and remove actuator (19) through access opening.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 9-14, safety
turnbuckle and reinstall all items removed for access.

9-7A. DISASSEMBLY. (Refer to figure 9-2A.)
a. Remove actuator in accordance with paragraph
9-7.
b. Disassemble actuator assembly (1) as illustrated
in Detail A as follows:
1. Remove chain guard (3) if not previously removed in step "e" of paragraph 9-7.
2. Using suitable punch and hammer, remove
roll pins (8) securing sprocket (5) to screw (9) and
remove sprocket from screw.
3. Unscrew threaded rod end (15) and remove
rod end from actuator.
4. Remove roll pins (10) securing bearings
(6 and 14) at the housing ends.
5. Lightly tap screw (9) toward the sprocket end
of housing, remove bearing (6) and collar (7).
6. Lightly tap screw (9) in the opposite direction from sprocket end, remove bearing (14), O-ring
(13) and collar (7).
7. It is not necessary to remove retaining
rings (11).
9-7B. CLEANING, INSPECTION AND REPAIR.
(Refer to figure 9-2A. )
a. DO NOT remove bearing (16) from threaded rod
end (15) unless replacement of bearing is necessary.
b. Clean all component parts, except bearing (16),
by washing Stoddard solvent or equivalent. Do not
clean sealed bearing (16).
c. Inspect all component parts for obvious indications of damage such as stripped threads, cracks,
deep nicks and dents.
d. Check bearings (6 and 14), screw (9) and threaded rod end (15) for excessive wear and scoring.
Dimensions of the parts are as follows:
BEARING (6)
INSIDE DIAMETER
0.370" MIN.
INSIDE DIAMETER
0. 373" MAX.
BEARING (14)
INSIDE DIAMETER
SMALL HOLE
0.248" MIN.
SMALL HOLE
0.253" MAX.
LARGE HOLE
0.373" MIN.
LARGE HOLE
0. 380" MAX.
THREADED ROD END (15)
OUTSIDE DIAMETER
(SHANK)
SCREW (9)
OUTSIDE DIAMETER

0.242" MIN.
0.246" MAX.
0. 367" MIN.
0.370" MAX.

h. DO NOT attempt to repair damaged or worn
parts of the actuator assembly. Discard all defective items and install new parts during reassembly.
9-7C. REASSEMBLY. (Refer to figure 9-2A.
a. Always discard the following items and install
new parts during reassembly.
1. Bearings (6 and 14)
2. Roll pins (8 and 10)
3. O-Ring(13)
4. Nuts (2).
b. During reassembly, lubricate collars (7), screw
(9) and threaded rod end (15) in accordance with
Section 2.
c. Press sprocket (5) into the end of screw (9),
align roll pin holes and install new roll pins (8).
d. Slip bearing (6) and collar (7) on screw (9) and
slide them down against sprocket (5).
e. Insert screw (9), with assembled parts, into
housing (12) until bearing (6) is flush with the end of
housing.
NOTE
When inserting screw (9) into housing (12),
locate the sprocket (5) at the end of housing
which is farther away from the groove for
retaining ring (11).
* The bearings (6 and 14) are not pre-drilled
and must be drilled on assembly. The roll
pins (10) are 1/16 inch in diameter, therefore, requiring a 1/16 (0.0625) inch drill.
f. With bearing (6) flush with end of housing (12),
carefully drill bearing so the drill will emerge from
the hole on the opposite side of housing (12). DO
NOT ENLARGE HOLES IN HOUSING.
g. Press new roll pins (10)into pin holes.
h. Insert collar (7), new O-ring (13) and bearing
(14) into opposite end of housing (12).
i. Complete steps "f" and "g" for bearing (14).
j. If a new bearing (16) is required, a new bearing
may be pressed into the boss. Be sure force bears
against the outer race of bearing.
k. Screw the threaded rod end (15) into screw (9).
1. Install retaining rings (11), if they were removed.
m. Test actuator assembly by rotating sprocket (5)
with fingers while holding threaded rod end (15).
The threaded rod end should travel in and out smoothly, with no indication of binding.
n. Reinstall actuator assembly in accordance with
paragraph 9-7.

NOTE
Relative linear movement between internal
threaded screw (9) and bearing (14) should
be 0.004 to 0.010 inch at room temperature.
e. Examine threaded rod end (15) and screw (9)
for damaged threads or dirt particles that may
impair smooth operation.
f. Check sprocket (5) for broken, chipped and/or
worn teeth.
g. Check bearing (16) for smoothness of operation,

9-7D. TRIM TAB FREE-PLAY INSPECTION.
a. Place elevators and trim tab in the neutral position.
b. Using moderate pressure, move the trim tab
trailing edge up and down by hand to check free-play.
c. A maximum of. 166",(total motion up and down)
measured at the trim tab trailing edge is permissible.
d. If the trim tab free-play is less than .166", the
system is within prescribed limits.
e. If the trim tab free-play is more than .166",
check the following items for looseness while moving
the trim tab up and down.
Change 1

9-2A

1. Check push-pull tube to trim tab horn assembly attachment for looseness.
2. Check push-pull tube to actuator assembly
threaded rod end attachment for looseness.
3. Check actuator assembly threaded rod end
for looseness in the actuator assembly with push-pull
tube disconnected.

SHOP NOTES:

9-2B

Change 1

f. If looseness is apparent while checking steps
e-1 and e-2, repair by installing new parts.
g. If looseness is apparent while checking step e-3,
refer to paragraphs 9-6 through 9-7C. Recheck trim
tab free-play.

REFER TO FIGURE 9-3
REFER TO
FIGURE 9-2
REFER TO FIGURE 9-4

11

1

..

Detail D

Detail A
_

1.
2.
3.
4.
5.

_

24
10

7.
8.
Detail

11

/

H

F

9.
15
11

Detail

H

10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.

Sprocket
Bearing
Bushing
Cable Guard
Spacer
Pulley
6.
Left Forward Cable
Cable Ends
Left Aft Cable
Right Aft Cable
Turnbuckle
Right Forward Cable
Travel Stop Block
Horn Assembly
Push-Pull Tube
Brace
Stabilizer Rear Spar
Support Bracket
Actuator
Sprocket
Chain Guard
Clamp
Chain

-

CRAFT SERIAL U20601701
CRAFT SERIAL U20601701
Detail D

ICAUTIONI

24. Bushing

MAINTAIN PROPER CONTROL
CABLE TENSION.

20

14
22
15

Figure 9-1.

CABLE
10 TO
ATURE
REFER

TENSION:
15 LBS (AT AVERAGE TEMPERFOR THE AREA.)
TO FIGURE 1-1 FOR TRAVEL.

Elevator Trim Tab Control System
Change

1

9-3

2

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

Right Elevator
Trim Tab
Hinge Half
Spacer
Foam Filler
Horn Assembly
Bushing
Bolt
Push-Pull Tube
Hinge Pin
Screw
Nutplate
Left Elevator

12
Detail B

9-8. TRIM TAB CONTROL WHEEL.
ure 9-3.)

Elevator Trim Tab Installation

(Refer to fig-

9-9. REMOVAL AND INSTALLATION.
a. Remove pedestal cover as outlined in paragraph
9-13.
b. Remove screws (8) and nuts (6) securing chain
guard (7) to pedestal structure (9).
c. Remove nut (4) securing Indicator (2) to pivot
stud (1). Retain washers (3) for reinstallation.
d. Loosen bolts (12) securing idler sprockets (11)
to pedestal structure (9), slide idler sprockets in
slotted holes and disengage chain (13) from sprockets.
e. Remove bolts (12) and remove chain guard (7)
using care not to bend indicator (2) or drop parts into
tunnel area.
f. Remove roller chain (13) from trim wheel sprocket and carefully slide wheel (5) from pivot stud (20).
g. Reverse the preceding steps for reinstallation.
Remove roller chain (13) slack by adjusting idler
sprockets (11) in slotted holes and reinstall all items
removed for access.
CABLES AND PULLEYS.

9-11. REMOVAL AND INSTALLATION.
a. FORWARD CABLE. (WITHOUT ELECTRIC
TRIM.) (Refer to figure 9-1. )
1. Peel back carpeting as necessary to expose
access plates in cabin and baggage areas and remove
plates.
9-4

Change I

Detail A

i

Figure 9-2.

9-10.

9

2. Remove safety wire, relieve cable tension
and disconnect turnbuckle (11).
3. Disconnect cable ends (8).
4. (Refer to figure 9-3.) Remove pedestal cover as outlined in paragraph 9-13.
5. Remove lower pedestal panel (19) and disengage roller chain (15) from drive sprocket assembly
(16).
6. Remove cable guards and pulleys as necessary to work cable free of aircraft.
NOTE
To ease routing of cable, a length of wire may
be attached to the end of cable before being
withdrawn from aircraft. Leave wire in place,
routed through structure; then attach the cable
being installed and pull cable into position.
7. Reverse the preceding steps for reinstallation.
8. After cable is routed in position, install pulleys and cable guards. Ensure cable is positione in
pulleygrooves before installing guards. Ensure roll-i
er chain (15) is positioned correctly over drive
sprocket (16).
9. Re-rig system in accordance with paragraph
9-14, safety turnbuckle (index 11, figure 9-1) and reinstall all items removed for access.
b. FORWARD CABLE. (WITH ELECTRIC TRIM.)
(THRU AIRCRAFT SERIALS P20600648 AND U20601700.) (Refer to figure 9-5.)

1.
2.
3.
4.
5.
6.
7.
8.
9.

Actuator Assembly
Nut
Chain Guard
Screw
Sprocket
Bearing
Collar
Pin
Screw

10. Pin

Detail

A

8

11. Retaining Ring
11. Housing
12.
Housing Ring
13. O-Ring
14. Bearing
15. Threaded Rod End
16. Bearing
17. Screw Assembly
18. Grease Zerk

NOTE
Disassembly, cleaning , inspection and
repair of tab actuator illustrated in Detail B is limited to replacement of guard
(3), sprocket (5), screw assembly (17),
zerk (18) and bearing (16). Other items
found defective will require actuator
assembly replacement as a unit. Lubrcate actuator in accordance with Section
2.

4

* NOTE
Used with electric
trim installation
BEGINNING WITH SERIAL U20602200

Figure 9-2A.

Elevator Trim Tab Actuator Assembly
Change 3

9-4A/(9-4B blank )

*

1

.
17

~~6--"t^cd~
\ |I

~*

Use as required to
hold indicator pivot
end away from trim
wheel (5).

A
It

/
^/^^fI~~~~

>^.N^

j16
\

10
7

s\

\

14
\

\^v

^

\

y ^^
"\ }

16~

~10

i\

\I

AdzA
\\i

1

4

_"~
~~~\

j~

Detail A

\/

91\~

r

aJ S\\

Figure 9-3.

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

Pivot Stud
Position Indicator
Washer
Nut
Trim Wheel
Nut
Chain Guard
Screw
Pedestal Structure
Bushing
Idler Sprocket

12.

Bolt

13.
14.
15.
16.
17.
18.

Roller Chain
Bolt
Roller Chain
Drive Sprocket
Lower Pedestal Panel
Pivot Stud

Elevator Trim Wheel Installation
Change

1

9-5

L

FWD

1.

With elevators in neutral, set trim tab to neutral (streamlined)

2.

Position stop block (1) against clevis on cable B and secure to cable B.

3.

Place inclinometer on trim tab and lower tab to degree specified in figure 1-1.

4.

Position stop block (2) against stop block (1) and secure to cable A.

5.

Raise trim tab to specified degree, place stop block (3) against stop block (2)
and secure to cable B.
Figure 9-4.

Elevator Trim Tab Travel Stop Adjustment

1. Peel back carpeting as necessary to expose
access plates in cabin and baggage areas and remove
plates.
-7.
2. Remove safety wire, relieve cable tension
and disconnect turnbuckle (6).
3. Disconnect cable ends (9) shown in Detail B
forward of the electric trim installation.
4. Complete steps 4 thru 9 of subparagraph "a."
c. FORWARD CABLE. (WITH ELECTRIC TRIM.)
(BEGINNING WITH AIRCRAFT SERIAL U20601701.)
(Refer to figure 9-6. )
1. Peel back carpeting as necessary to expose
access plates in cabin and baggage areas and remove
plates.
2. Remove safety wire, relieve cable tension
and disconnect turnbuckle (28).
3. Disconnect clamps and keepers (36) from left
forward cable (30).
4. Disconnect cables (29 and 30) at cable ends.
5. Complete steps 4 thru 9 of subparagraph "a."
d. AFT CABLE. (WITHOUT ELECTRIC TRIM.)
(Refer to figure 9-1. )
1. Remove rear baggage compartment wall.
2. Remove safety wire, relieve cable tension
and disconnect turnbuckle (11).
JCAUTION 1
Position a support stand under tail tie-down
ring to prevent tailcone from dropping while
working inside.
3. Disconnect cable ends (8).
4. Remove travel stop blocks (13).
5. Remove access plate beneath trim tab actuator (19) and remove chain guard (21).

9-6

Change 3

6. Disengage roller chain (23) from actuator
sprocket (20).
Remove cable guards and pulleys as necessary to work cable free of aircraft.
NOTE
To ease routing of cable, a length of wire
may be attached to the end of cable before
being withdrawn from aircraft. Leave
wire in place, routed through structure;
then attach the cable being installed and
pull cable into position.
8.

Reverse the preceding steps for reinstalla-

tion.
9. After cable is routed in position, install pulleys and cable guards. Ensure cable is positioned
in pulley grooves before installing guards. Ensure
roller chain (23) is positioned correctly over actuator sprocket (20). Ensure bushing (24) is positioned
in stop blocks (13).
10. Re-rig system in accordance with paragraph
9-14, safety turnbuckle (11) and reinstall all items
removed for access.
e. AFT CABLE (WITH ELECTRIC TRIM.) (THRU
AIRCRAFT SERIALS P2060064& AND 32qQI0601100.
(Refer to figure 9-5.)
1. Complete step 1 of subparagraph "d."
2. Remove safety wire, relieve cable tension
and disconnect turnbuckle (6).
{CAUTION

I

Position a support stand under tail tie-down
ring to prevent tailcone from dropping while
working inside.

3. Disconnect cable ends (9) shown in Detail B
aft of the electric trim installation.
4. Remove travel stop blocks (3).
5. (Refer to figure 9-1.) Complete steps 6 thru
11 of subparagraph "d."
f. AFT CABLE. (WITH ELECTRIC TRIM.)
(BEGINNING WITH AIRCRAFT SERIAL U20601701.)
(Refer to figure 9-6.)
1. Complete steps I and 2 of subparagraph "d."
2. Remove safety wire, relieve cable tension
and disconnect turnbuckle (28).

CAUTION
Position a support stand under tail tie-down
ring to prevent tailcone from dropping while
working inside.
3. Disconnect cables (29 and 30) at cable ends.
4. Remove travel stop blocks (2).
5. (Refer to figure 9-1.) Complete steps 6 thru
11 of subparagraph "d."
9-12.

PEDESTAL COVER.

9-13. REMOVAL AND INSTALLATION.
a. Turn fuel selector valve to OFF position and
drain fuel from strainer and lines.
b. Remove knurled nut from engine primer if installed and pull plunger from primer body. Protect
primer from dirt.
handle
and placard.
d.c. Remove
Remove fuel
cowl selector
flap
handle/knob.
breaker
nut and
trim circuit
e. Remove
electric
microphone
mounting
bracket,
if installed.
f. Fold carpet back as necessary and remove
screws securing cover to floor and pedestal.
g. Disconnect electrical wiring to pedestal lights.
h. Carefully work cover from pedestal to prevent

damage.
i.

Reverse the preceding steps for reinstallation.

9-14. RIGGING - STANDARD TRIM SYSTEM.
(Refer to figure 9-1.)
CAUTION
Position a support stand under tail tie-down
ring to prevent tailcone from dropping while
working inside.
a. Remove rear baggage compartment wall and
access plates as necessary.
b. Loosen travel stop blocks (13) on trim tab
cables (9 and 10).
c. Disconnect push-pull tube (15) from actuator
(19).
d. Check cable tension for 10-15 pounds and readjust turnbuckle (11), if necessary.
NOTE
If roller chains and/or cables are being installed, permit actuator screw to rotate
freely as roller chains and cables are connected. Adjust cable tension and safety
turnbuckle (11).

e. (Refer to figure 9-3.) Rotate trim control wheel
(5) full forward (nose down). Ensure pointer (2) does
not restrict wheel movement. If necessary to reposition pointer, proceed as follows:
1. Remove pedestal cover as outlined in paragraph 9-13.
2. Loosen nut (6) at trim wheel pivot stud (20).
3. Loosen screws (8) securing chain guard (7)
far enough that trim wheel (5) can be moved approximately 1/8 inch, then reposition pointer (2) using a
thin screwdriver to pry trailing leg of pointer out of
groove in trim wheel. Reposition pointer as required.
4. Tighten nut (6) and screws (8), but do not reinstall pedestal cover until rigging is complete.
NOTE
Full forward (nose down) position of trim
wheel is where further movement is prevented by the roller chain or cable ends
contacting sprockets or pulleys.
f. With elevator and trim tab both in neutral
(streamlined), mount an inclinometer on trim tab
and set to 0 ° . Disregard counterweight areas of
elevators when streamlining. These areas are contoured so they will be approximately 3° down at
cruising speed.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
g. Rotate actuator screw in or out as required to
place trim tab up with a maximum of 2' overtravel,
with actuator screw connected to push-pull tube (index 15, figure 9-1).
h. Rotate trim wheel to position trim tab up and
down, readjusting actuator screw as required to
obtain overtravel in both directions.
i. Position stop blocks and adjust as illustrated in
figure 9-4 to degree of trim tab travel specified in
figure 1-1.
j. Install pedestal cover and adjust trim tab pointer
(2) as follows:
1. Rotate trim control wheel (5) to place tab at
10 ° up position.
2. Locate the pointer (2) at the "TAKE-OFF"
triangle as viewed from the pilot seat. (Refer to
step "e," and reposition pointer if necessary.)
3. Bend pointer (2) as required to clear pedestal cover. (Pointer must NOT rub against pedestal
cover or clear cover more than .125 inch maximum.)
k. Safety Turnbuckle and reinstall all items
removed in step "a".

WARNING
Be sure trim tab moves in correct direction
when operated by trim control wheel. Nose
down trim corresponds to tab up position.

Change 1

9-7

9-15. ELECTRIC TRIM ASSIST INSTALLATION.
(Refer to figure 9-5, 9-6 and 9-7.)

disengage switch, the other switch operating electric
trim assist. The electric trim circuit breaker is
mounted on pedestal cover, the electrical wiring is
routed thru cabin and fuselage to Sta. 209.00 then
routed UP thru elevator to voltage regulator and
drive assembly. The drive assembly includes a gear
motor and two sprockets that operates a chain driven,
solenoid-operated, adjustable clutch. The actuator
assembly has dual sprockets. The manual trim tab
UP cable connects to the actuator around the AFT
sprocket. The drive assembly connects to the actuator by a chain around the FWD sprocket. When the
clutch is not energized, the drive drum "free wheels"
and has no effect on manual operation.

9-16 DESCRIPTION. AIRCRAFT SERIALS P20600648 THRU U20602199. The electric trim assist
is operated by a control wheel-mounted switch. The
servo unit includes a motor and a chain driven,
solenoid-operated, adjustable clutch. The trim tab
UP cable enters the servo housing and double wraps
around a drive drum. When the clutch is not energized, the drive drum "free wheels" and has no
effect on manual operation. AIRCRAFT BEGINNING
WITH SERIAL U20602200 (Refer to figure 9-7.) The
electric trim assist is operated by two switches
mounted on control wheel one switch operating the
9-17.

TROUBLE SHOOTING

TROUBLE
SYSTEM INOPERATIVE.

TRIM MOTOR OPERATING TRIM TAB FAILS TO MOVE.

PROBABLE CAUSE
Circuit breaker out.

Check visually.

Defective circuit breaker.

Check continuity.
breaker.

Replace defective

Defective wiring.

Check continuity.

Repair wiring.

Defective trim switch.

Check continuity.
switch.

Replace defective

Defective trim motor.

Remove and bench test.
defective motor.

Defective clutch solenoid.

Check continuity.
solenoid.

Improperly adjusted clutch
tension.

Check and adjust spanner nuts
for proper tension.

Disconnected or broken
cable.

Operate manual trim wheel.
Connect or replace cable.

Defective actuator.

Check actuator operation.
Replace actuator.

9-18. REMOVAL AND INSTALLATION.
a. THRU AIRCRAFT SERIALS P20600648 AND
U20601700. (Refer to figure 9-5.)
1. Remove aft baggage compartment wall.
2. Remove safety wire and relieve cable tension
at turnbuckle (6).

Position a support stand under tail tie-down
ring to prevent the tailcone from dropping
while working inside.
3. Disconnect left center cable (12) at both cable
ends (9).
4. Disconnect electrical wiring to servo unit.
5. Remove mounting bolts (10) and remove unit
from aircraft.
9-8

Change 1

REMEDY
Reset breaker.

Replace

Replace

6. Reverse the preceding steps for reinstallation. Rig system in accordance with paragraph 9-21,
safety turnbuckle (6) and reinstall all items removed
for access.
b. BEGINNING WITH AIRCRAFT SERIAL U20601701
THRU U20602199 (Refer to figure 9-6.)
1. Remove aft baggage compartment wall.
2. Disconnect electric trim assist cable (35) at
both
ends by removing clamps and keepers (36).
3. Remove cable guard (25) from bracket (26).
4. Disconnect electrical wiring to servo unit.
5. Remove mounting bolts (22) and remove unit
from aircraft.
6. Reverse the preceding steps for reinstallation. Check system rigging in accordance with paragraph 9-21 and re-rig, if necessary.

REFER TO FIGURE 9-4

REFER TO FIGURE 9-2

A

Detail

-D

1.
2.
3.
4.
5.

6. Turnbuckle
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.

B

Circuit Breaker
Trim Switch
Travel Stop Block
Trim Tab
Cable (Right Forward)
Cable (Rght Aft)
Cable (Left Aft)
Cable Ends
Mounting Bolt
Pulley
Cable (Left Center)
Cable Guard
Grommet
Cover
Bearing
Housing
Roller Chain
Sprocket
Motor Support
Motor
Motor Cover
Clutch Cover
Bushing
Spanner Nut
Washer
Friction Washer
Drive Drum
Shalt Assembly
Solenoid Clutch
Support Structure
Stiffener
Spacer
Voltage Regulator\
Connector
Spring Scale
-

The clutch setting is0
25 + 2 - O lb in.

E

*
/
/

'

,

\

'<

32

2312
24

13
\>

25
2t
/

i
i

\3
2
f-q
'

\

'

1
14

.
14
.o

/

15

21
15
IS
17

Detail

11
2
fHRU AIRCRAFT SERIALS
P20600648 AND U20601700

*Safety wire these items.
Remainder of the elevator trim systern is illustrated in figure 9-1.
Refer to Section 2 for lubrication requirements.
Figure 9-5.

Detail

ICA

t

CAUTIONI

MAINTAIN PROPER CONTROL
CABLE TENSION.
CABLE TENSION:
20 LBS + 5 - 0 LBS (AT AVERAGE TEMPERATURE FOR THE AREA.)
REFER TO FIGURE 1-1 FOR TRAVEL.

Electric Elevator Trim System thru P20600648 & U20601700 (Sheet 1 of 2)
Change 1

9-9

Detail D
* Spacer (33) replaces the pulley
normally installed in the standard system when the electric
trim system is installed in the
aircraft.

* Support (31) is rotated 90' to

expose voltage regulator (34).

NOTE
Beginning with aircraft serial
U20601588 a 24 volt electrical
system may be installed.

NOTE
*With an external power source supplying
27.5 volts to the aircraft, adjust the
voltage regulator (34) to 10 volts output
in both directions.

NOTE
*

Figure 9-5.
9-10

Change 1

Detail E applies only to aircraft
serials U20601588 thru U20601700
when equipped with a 24 volt electrical system and an electric trim
system. Detail E does not apply to
12 volt systems equipped with
electric trim.

Electric Elevator Trim System thru P20600648 & U20601700 (Sheet 2 of 2)

1.
2.
3.
4.
5.

Trim Tab
Travel Stop Block
Trim Switch
Circuit Breaker
Clutch Cover

REFER TO FIGURE 9-2
REFER TO FIGURE 9-4

2

3

6. Bearing
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
37.
38.

Bushing
Spanner Nut
Washer
Friction Washer
Drive Drum
Shaft Assembly
Solenoid Clutch
Sprocket
Roller Chain
Grommet
Motor Cover
Motor
Motor Support
Cover
Housing
Mounting Bolt
Support Assembly
Pulley
Cable Guard
Bracket
Cable (Right Aft)
Turnbuckle
Cable (Left Aft)
Cable (Left Forward)
Cable (Right Forward)
Spacer
Voltage Regulator
Swaged Ball
Assist Cable
Clamp and Keeper
Connector
Spring Scale

4

B
*
5
/ 6 7

'
..

.
8

|17

10

19
9

\4
12
/ I.
4 1
W
Thru aircraft serial U20601748, the clutch setting
Is 20 ± 1 lb in.

4
15

Beginning with aircraft
serial
al- U20601749, the
.
clutch setting is 30 + 0
- 2 lb in.
20

\

21
6

* Safety wire these items.

18

2
I

DetailC

Remainder of elevator trim sys-J/

,

.

tem is illustrated in figure 9-1.

\

j
'

23

Refer to Section 2 for lubrication requirements,

r
,^ ,A

B':
2>^y

30
Detail

A

\0^

^^/

<
^\

8~~~~25

A

BEGINNING WITH AIRCRAFT SERIAL U20601701 THRU U20602199

Figure 9-6.

Electric Elevator Trim System Beginning U20601701 Thru U20602199 (Sheet 1 of 2 )
Change 1

9-11

* Support (23) is rotated 90* to
expose voltage regulator (33).
*33

34
/

11

21

37
23
Detail D

.
//

35

26

VIEWA A

AIRCRAFT C/L
33

21

NOTES
Assist cable (35) must be wrapped around
drive drum (11) so that the threaded end

/

.

-24

of the assist cable exits the housing on the
/23

same side where the motor end is visable.

* With an external power source supplying
13.75 volts to the aircraft. (when equipped

with 12 volt electrial system ) or 27.5
volt when (aircraft equipped with 24 volt
system ) adjust the voltage regulator (33)
to 10 volts output in both directions.

_

/
/
25

26

VIEW
BB

Figure 9-6.

9-12 Change 1

Electric Elevator Trim System Beginning U20601701 (Sheet 2 of 2)

24

1.
2.

Trim Tab
Voltage Regulator

3. Trim Switch
4.
5.
6.
7.
8.
9.
10.
11.

Disengage Switch
Circuit Breaker
Push-Pull Tube
Brace
Mounting Bracket Assembly
Actuator Assembly
Mounting Plate - FWD.
Sprocket Guard

12. Washer
13. Shaft
14. Sprocket
15. Chain Assembly
16. Washer Assembly

17.

Spring Washer

18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
37.
38.
39.

Washer
Nut
Shaft Assembly
Clutch
Cover Assembly
Rub Strip
Sprocket
Bushing
Chain
Gear Motor
Mounting Plate
Sprocket
Pin
Housing Assembly
Cover
Mounting Plate Assembly
Drive Assembly
CTR1 Adjustment
CTR2 Adjustment
Connector
Cover
Cover
SAFETY WIRE THESE ITEMS

Figure 9-7.

Electric Elevator Trim System Beginning U20602200 (Sheet 1 of 2)
Change 1

9-12A

14
23

222

20

3

3

\\>;\a

1

etC2
29
30

32

Detail C

37

CONNECTOR

I

35

^4'

;

NOTE
.i*^^'t
o0\
J^j^ 1, 3,2-

*

Used only on aircraft equipped with
24 volt electrical system.

36,Detail

Figure 9-7.
9-12B

Change 3

D

Electric Elevator Trim System Beginning U20602200 (Sheet 2 of 2)

e. AIRCRAFT WITH OPTIONAL ELECTRIC
TRIM ASSIST INSTALLATION BEGINNING WITH
SERIAL U20602200 (Refer to figure 9-7.)
1. Remove access plate below actuator and
rnvprs (38) & (39).
2. Disconnect electric trim assist cable (37) and
three Mate-N-Lok connectors on drive assembly.
Remove bolt and nut from ground wire thru rib.
3. Remove sprocket guard (11) from actuator
body,
4. Remove mounting bolts from voltage reulator
(2) and drive assembly (34) actuator (9) and remove
units from aircraft.
5. Reverse the preceding steps for reinstallation. Check system rigging in accordance with paragraph 9-21 and safety wire turnbuckle if re-rigging
is necessary.
9-19. CLUTCH ADJUSTMENT.
a. THRU AIRCRAFT SERIALS P20600648 AND
U20601700. (Refer to figure 9-5.)
1. Remove aft baggage compartment wall.
2. Remove safety wire and relieve cable tension
at turnbuckle (6).
3. Disconnect left center cable (12) at both
cable ends (9).
4. Disconnect electrical power to the motor
assembly (21) by unplugging the connector installed
in the RED wire leading to the motor assembly.
NOTE
Step 4 isolates the motor assembly from
the remainder of the electric trim system
so it cannot be engaged during clutch adjustment.

NOTE
Spanner nuts (25) may be loosened or tightened
with a suitable hammer and punch.
12. Repeat steps 10 and 11 until tension is in
accordance with step 10, then tighten outside spanner
nut against inside nut.
13. Connect electrical wiring to motor assembly
which was removed in step 4, re-rig trim system in
accordance with paragraphs 9-14 and 9-21 and reinstall all items removed for access.
b. BEGINNING WITH AIRCRAFT SERIAL U20601701.
THRU U20602199 ( Refer to figure 9-6
1. Remove aft baggage compartment wall.
2. Disconnect assist cable (35) at both ends by
removing clamps and keepers (36).
3. Disconnect electrical power to the motor
assembly (18) by unplugging the connector installed
in the RED wire leading to the motor assembly.
NOTE
Step 3 isolates the motor assembly from the
remainder of the electric trim system so it
cannot be engaged during clutch adjustment.
4. Remove screws securing cover (20) to housing (21) and slide the cover down over electrical wlring far enough to expose the clutch assembly.
5. Ensure the electric trim circuit breaker on
the pedestal cover is pushed IN and place master
switch in the ON position.
6. Operate control wheel-mounted switch UP or
DOWN to energize the solenoid clutch (13).
7. Attach the spring scale (38) to the assist
cable (35) and pull scale slowly until slippage is
noticed.
Slippage should occur between 33. 86 to 37.25 bs on
12 and 24 volt aircraft systems.

5. Remove screws securing cover (15) to housing
(17) and slide the cover down over electrical wiring
far enough to expose the clutch assembly.
6. Ensure the electric trim circuit breaker on
. Repeat steps 6 and 7 several tlmes to break
the initial friction of the clutch, making sure that
the pedestal cover is pushed IN and place master
cable (35) is re-wound on drive drum (11) after each
switch in the ON position,
slippage test.
7. Operate control wheel-mounted switch UP or
DOWN to energize the solenoid clutch (30).
9. Repeat steps 7 and 8 very slowly, carefully
8. Attach the spring scale (38) to the left center
watching the indicator on the spring scale (38).
10. If tension is not within tolerance, loosen
cable (12) and pull scale slowly until slippage is
OUTSIDE spanner nut (8) which act as a lock.
~~~~~~~~noticed.
~Tighten
INSIDE spanner nut to increase clutch ten9. Repeat steps 7 and 8 several times to breakn
nut to crease clutch ten
sion and loosen nut to decrease clutch tension.
the initial friction of the clutch, making sure that
cable (12) is re-wound on drive drum (28) after each
NOTE
slippage test.
10. Repeat steps 7 and 8 very slowly, carefully
S
n
( m
b loos
or i
may be loosened or tightened
nuts (81 ammer
Spamer
watching the indicator on the spring scale (38).
it aa suitable
and
with
suitable hammer
and pun.
punch.
Slippage should occur between 28.22 to 30.47 lbs on
12 volt aircraft systems and between 21.44 to 23.70.
Repeat steps 9 and 10 until tension is in
Repeat steps 9 and l0 until tension Is in
,,11.
fIbs on 24 volt aircraft systems.
with step 9, then tighten outside spanner
accordance
loosen
tolerance,
within
not
is
If
tension
11.
11. If tension is not within tolerance, loosen
nut aainst inside nut.
nut against inside nut.
OUTSIDE spanner
spanner nut
nut (25)
(25) which
which acts
acts as
as aa lock.
lock
nu 12.
ain Connect
i
nut.
electrical
wiring to motor assembly
OUTihSIDE
Tighten
NSDEspanner
nut to
crease
inde clutch tension
which was removed in step 3, re-rig trim system in
sion and loosen nut to decrease clutch tension.
accordance with paragraphs 9-14 and 9-21 and reinaccordance with paragraphs 9-14 and 9-21 and reinstall all items removed for access.

Change 1

9-13

I

BEGINNING WITH AIRCRAFT SERIAL U20602200
(Refer to figure 9-7.)
1. Remove access plate below actuator and
covers (38) & (39).
2. Remove safety wire and relieve cable tension
and chain tension at turnbuckles.
3. Disconnect electric motor by unplugging the
three Mate-N-Lok connectors leading to the motor
assembly.
4. Remove mounting bolts from drive assembly.
It is necessary to remove from elevator to make the
necessary adjustments to clutch.
NOTE
Step 3 isolates the motor assembly from the
remainder of the electric trim system so it
cannot be engaged during clutch adjustment.
5. Remove screws securing covers (23) and (22)
to housing (31) and slide the cover down over electrical wiring far enough to expose the clutch assembly.
6. Ensure the electric trim circuit breaker on
the pedestal cover is pushed in and place master
switch in the ON position.
7. Operate control wheel- mounted switch UP or
DOWN to energize the solenoid clutch (21).

c. Disconnect the electrical power leads to the
motor by unplugging the connectors installed in the
RED and BLACK wires leading to the motor assembly.
d. Connect one lead of a dc voltmeter capable of
measuring the aircraft voltage to either the RED or
BLACK wire leading to the motor and the other voltmeter lead to a good aircraft ground.
e. Operate the electric trim switch to the NOSE UP
and NOSE DOWN positions and check voltage present
at the RED and BLACK wires.
f. Adjust CTR 1 and CTR 2 adjustment screws on
the voltage regulator counterclockwise (CCW), then
slowly turn adjustment screws clockwise (CW) until
a 10 volt output is obtained for both (RED and BLACK)
leads.
g. Remove voltmeter and reconnect the motor assembly power leads. Be sure to connect RED to RED
and BLACK to BLACK when reconnecting leads.
h. Check trim system for proper operation and reinstall all items removed for access.

9-20A. DUAL VOLTAGE REGULATOR ADJUSTMENT. (24 VOLT SYSTEM ONLY BEGINNING WITH
U20602200)
(Refer to figure 9-7.)
a. Remove access cover (39).
8.
Attach the spring scale (Index (38) in Figure
b. Connect an external power source of 13.75 volts
9-6 to chain and pull scale slowly until slippage is
(aircraft equipped with 12 volt electrical systems) or
noticed.
27.5 volts (aircraft equipeed with 24 volt electrical
9. Repeat
Steps 7 &8 several times to break
systems) dc
continuous to the aircraft electrical sys-

the initial friction of the clutch.
tem or if an external power supply is not available,
10. Repeat Steps 8 and 9 very slowly, carefully
run the aircraft engine at approximately 1000 RPM to
watching the indicator on the spring scale. Slippage
maintain the normal operating aircraft voltage.
should occur between 29.1 to 32.9 lbs. on 12 and 24
volt aircraft systems.
c. Disconnect the electrical power leads to the
11. IF tension is not within tolerance, loosen
motor by unplugging the connectors installed in the
OUTSIDE spanner nut (19) which acts as a lock.
RED and BLACK wire leading to the motor assembly.
Tighten INSIDE spanner nut to increase clutch tend. Connect one lead of a dc voltmeter capable of
sion and loosen nut to decrease clutch tension.
measuring the aircraft voltage to either the RED or
BLACK wire leading to the motor and the other voltNOTE
meter lead to a good aircraft ground.
e. Operate the electric trim switch to the Nose UP
Spanner nut (19) may be loosened or tightened
and Nose DOWN positions and check voltage present
with a suitable hammer and punch.
at the RED and BLACK wires.
f. Adjust CTR 1 and CTR 2 adjustment screws on
12. Repeat Steps 10 and 11 until tension is in
the voltage regulator counterclockwise (CCW). then
accordance with 10. then tighten outside spanner nut
slowly turn adjustment screws clockwise (CW) until
against inside nut.
a 13.5 volt output is obtained for both (RED and
13. Connect electrical wiring to motor assembly
BLACK ) leads.
which was removed in Step 3, re-rig trim system in
g. Remove voltmeter and reconnect the motor asaccordance with paragraphs 9-14 and 9-21 and resembly power leads. Be sure to connect RED to RED
install all items removed for access.
and BLACK to BLACK when reconnecting leads.
h. Check to see if full "NOSE UP" to full "NOSE
9-20. DUAL VOLTAGE REGULATOR ADJUSTMENT.
DOWN" and full "NOSE DOWN" to full "NOSE UP"
(Beginning with aircraft serials U20601588 (24 volt
cycle time is 32 + or -3 seconds.
systems only) and U20601701 (12 volt and 24 volt sysReadjust voltage regulator as required to obtain
tems.)
32±3 seconds cycle time.
a. Remove the aft baggage compartment wall.
j. Check trim system for proper operation and reb. Connect an external power source of 13.75 volts
install all items removed for access.
(aircraft equipped with 12 volt electrical systems)
or 27. 5 volts (aircraft equipped with 24 volt electriCAUTION
cal systems) dc continuous to the aircraft electrical
The
trim motor should be allowed to cool
system, or if an external power supply is not availbetween voltage regulator adjustments for
able, run the aircraft engine at approximately 1000
approximately 5 minutes if several actuarpm to maintain the normal operating aircraft volttions of the motor becomes necessary durage.
ing adjustment.
9-14

Change 3

9-21. RIGGING - ELECTRIC TRIM ASSIST.
a. THRU AIRCRAFT SERIALS P20600648 AND
U20601700. (Refer to figure 9-5.)
1. The standard manual elevator trim control
system MUST be rigged in accordance with paragraph 9-14 prior to rigging the electric trim assist.
2. Remove rear compartment baggage wall.
3. Remove safety wire and adjust turnbuckle
(6) to increase trim system cable tension from 10
to 15 lbs to 20+5-0 lbs.
4. Recheck trim tab travel with an inclinometer
for degree of travel specified in figure 1-1. safety
turnbuckle (6) and reinstall all items removed for
access.
b. AIRCRAFT SERIALS U20601701 THRU
U20601748. (Refer to figure 9-6.)
1. Complete steps 1 and 2 of subparagraph "a."
2. Disconnect assist cable (35) at both ends by
removing clamps and keepers (36).
3. Remove safety wire and adjust turnbuckle
(28) to increase trim system cable tension from 10
to 15 lbs to 20+5-0 lbs.
4. Rotate trim control wheel to place trim tab
in the approximate mid-travel position (10 ° up).
5. Index the swaged ball (34) to the top of drive

drum (11).
6. Connect assist cable (35) to left forward
cable (30) and adjust the assist cable to 25+5-0
pounds tension.
7. Recheck trim tab travel with an inclinometer
for degree of travel specified in figure 1-1, safety
turnbuckle (28) and reinstall all items removed for
access.
c. AIRCRAFT SERIAL U20601749 THRU U20602199
(Refer to figure 9-6. )
1. Complete steps 1 thru 5 of subparagraph "b.
2. Connect assist cable (35) to left forward
cable (30) and adjust the assist cable to 10+5-0
pounds tension.
d. BEGINNING WITH AIRCRAFT SERIAL U20602200 (Refer to figure 9-7.)
1. Complete steps 1 and 2 of subparagraph "a"
2. Rig electric trim drive chain as follows:
a. Move elevator trim tab to full "NOSE UP"
position.
b. Locate NAS288 terminal on upper side of.
chain at a point 0. 75 inches from drive
assembly housing.
c. Adjust AN155 barrel until chain deflection
between Sprockets is approximatley 0. 25 inch.
d. Resafety turnbuckle and reinstall all items
removed for access.

SHOP NOTES:

Change 3

9-15/(9-16 blank)

SECTION 10
RUDDER CONTROL SYSTEM

Page

TABLE OF CONTENTS

Removal and Installation ....
. . ..
Repair . . . ...
Cables and Pulleys ......
Removal and Installation ......
...........
Rigging.

10-1
RUDDER CONTROL SYSTEM ........
. 10-1
. .
...
. . ..
Description ..
10-1
Trouble Shooting ............
10-9
Rudder Pedal Assembly .........
10-9
Removal and Installation ......
Rudder . . . . . . . . . . . . . . . . 10-9

10-1. RUDDER CONTROL SYSTEM.
ure 10-1.)

(Refer to fig-

10-2. DESCRIPTION. Rudder control is maintained
through use of conventional rudder pedals which also
control nose wheel steering. The system is com-

10-3.

. ..

. 10-9
. 10-9
10-9
10-9
10-9

prised of the rudder pedals installation, cables and
pulleys, all of which link the pedals to the rudder
and nose wheel steering. When dual controls are
installed, stowable rudder pedals are provided at
the copilot's position.

TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 10-11.

TROUBLE
RUDDER DOES NOT RESPOND
TO PEDAL MOVEMENT.

PROBABLE CAUSE
Broken or disconnected cables.

REMEDY
Open access plates and check
visually. Connect or replace
cable s.

Change 1

10-1

10-3.

TROUBLE SHOOTING (Cont).
TROUBLE

BINDING OR JUMPY MOVEMENT OF RUDDER PEDALS.

PROBABLE CAUSE

REMEDY

Cables too tight.

Refer to figure 10-1 for cable
tension. Rig system in accordance with paragraph 10-11.

Cables not riding properly on
pulleys.

Open access plates and check
visually. Route cables correctly over pulleys.

Binding, broken or defective
pulleys or cable guards.

Open access plates and check
visually. Replace defective
pulleys and install guards
properly.

Pedal bars need lubrication.

Refer to Section 2.

Defective rudder bar bearings.

If lubrication fails to eliminate
binding. Replace bearing blocks.

Defective rudder hinge bushings.

Check visually.
bushings.

Clevis bolts too tight.

Check and readjust bolts to
eliminate binding.

Steering rods improperly
adjusted.

Rig system in accordance with
paragraph 10-11.

LOST MOTION BETWEEN
RUDDER PEDALS AND
RUDDER.

Insufficient cable tension.

Refer to figure 10- for cable
tension. Rig system in accordance with paragraph 10-11.

INCORRECT RUDDER TRAVEL.

Incorrect rigging.

Rig in accordance with paragraph
10-11.

STOWABLE PEDALS DO
NOT DISENGAGE.

Broken or defective control.

Disengage control and check
manually. Replace control.

STOWABLE PEDALS DO
NOT STOW.

Defective cover, catch or
latch pin.

Check visually.
parts.

STOWABLE PEDALS DO
NOT RE-ENGAGE.

Binding control.

Check control operation.
or replace control.

Misaligned or bent mechanism.

Check visually. Repair or replace
defective parts.

10-2

Replace defective

Replace defective

Repair

A

REFER TO FIGURE 10-4

REFER TO FIGURE 10-2

D

"'

THRU AIRCRAFT
SERIAL U20601905

.D/

77

H

A

A

^; pDetail
:

REFER TO
SECTION 11
FOR RUDDER
TRIM CONTROL
SYSTEM.

*<-ll

DetailD
Detail4
~ .

3

"'^

...

"/~

_

8.
10.
n
11.

~13.

Firewall
Shock-Mount

Clamp
Support

/

2

RIGHT EXHAUST

/

/

^~ji

/

~12.

/

Muffler
Clamp Hall
Tailpipe
Shroud
Cabin Heat Outlet

4

LEFT EXHAUST

34

13

',

THRU AIRCRAFT

Detail A
BEGINNING WITH AIR -

SERIAL U20601668

CRAFT SERIAL U20601669

Detail A

Figure 12-11.
12-34

Change 1

Exhaust System

12-92.

STARTER MOTOR.

12-93. REMOVAL AND INSTALLATION.
a. Remove engine cowling in accordance with paragraph 12-3.

CAUTION
When disconnecting starter electrical cable,
do not permit terminal bolt to rotate. Rotation of the bolt could break the conductor
between bolt and field coils causing the
starter to be inoperative.
b. Disconnect battery cables and insulate as a
safety precaution.
c. Disconnect electrical cable at starter motor.
d. Remove nuts and washers securing motor to
starter adapter and remove motor. Refer to engine
manufacturer's overhaul manual for adapter removal.
e. Reverse the preceding steps for reinstallation.
Install a new O-ring seal on motor, then install motor.
Be sure motor drive engages with the adapter drive
when installing.
12-94.

EXHAUST SYSTEM.

12-95. DESCRIPTION. The exhaust system consists
of two exhaust stack assemblies, for the left and right
bank of cylinders. Each cylinder has a riser pipe attached to the exhaust port. The three risers at each
bank of cylinders are joined together into a collector
pipe forming an exhaust stack assembly. The center
riser on each bank is detachable. but the front and aft
risers are welded to the collector pipe. The left muffler is enclosed in a shroud which captures exhaust
heat which is used to heat the cabin.
12-96. REMOVAL AND INSTALLATION. (Refer to
figure 12-11. )
a. Remove engine cowling in accordance with paragraph 12-3.
b. Disconnect ducts from heater shroud on left muffler assembly.
c. Disconnect tailpipe braces from shock-mounts at
firewall brackets.
d. Remove nuts, springs and bolts attaching tailpipe
and muffler to collector pipe and remove muffler and
tailpipe assemblies.
e. Remove nuts attaching exhaust stack assemblies
to the cylinders and remove exhaust stacks and gas-ketsand
f.t Reverse
Reverse the
the preceding
preceding steps
steps for
for reinstallation.
reinstallation.
Install
a new
o coppe
ggasket
e b
Install a new copper-asbestos
betweenn eeach
n eh
g
t b
a new ocopprriser and its mounting pad on each cylinder, regardless of apparent condition of those removed. Torque
exhaust stack nuts at cylinders to 100- 110 poundex
hasstcnustclnest1010puding
inches.
12-97 INSPECTION.
Since exhaust systems of this
type are subject to burning, cracking and general
deterioration from alternate thermal stresses and
vibrations, inspection is important and should be
accomplished every 100 hours of operation. Also, a
thorough
thorough inspection
inspection of
of the
the engine
engine exhaust
exhaust syst
system
should be made to detect cracks causing leaks which
could result in loss of engine power. To inspect the
engine exhaust system, proceed as follows:

a. Remove engine cowling as required so that ALL
surfaces of the exhaust assemblies can be visually
inspected.
NOTE
Especially check the areas adjacent to welds
and slip joints. Look for gas deposits in surrounding areas, indicating that exhaust gases
are escaping through a crack or hole or
around the slip joints.
b. After visual inspection, an air leak check should
be made on the exhaust system as follows:
1. Attach the pressure side of an industrial
vacuum cleaner to the tailpipe opening, using a rubber plug to effect a seal as required.
NOTE
The inside of the vacuum cleaner hose should
be free of any contamination that might be
blown into the engine exhaust system.
2. With vacuum cleaner operating, all joints
in the exhaust system may be checked manually by
feel, or by using a soap and water solution and
watching for bubbles. Forming of bubbles is considered acceptable, If bubbles are blown away
system is not considered acceptable.
c. Where a surface is not accessible for a visual
inspection, or for a more positive test, the following
procedure is recommended.
1. Remove exhaust stack assemblies.
2. Use rubber expansion plugs to seal openings.
3. Using a manometer or gage, apply approximately 1-1/2 psi (3 inches of mercury) air pressure
while each stack assembly is submerged in water.
Any leaks will appear as bubbles and can be readily
detected.
4. It is recommended that exhaust stacks found
defective be replaced before the next night.
d. After installation of exhaust system components
perform the inspection in step "b" of this paragraph
to ascertain that system is acceptable
12-98. EXTREME WEATHER MAINTENANCE.
12-99.
be me

COLD WEATHER. Cold weather starting will
easier by the installation of an oil dilution
a ground service
system, an engine primer system and a ground service
receptacle. The primer system is manually-operated
from the cabin. Fuel is supplied by a line from the
fuel strainer to the plunger. Operating the primer
fl
s
r to te p
. O
t
forces fuel to the engine. With an external power refoes
fel o he engin.
Wit an eterna
er reThe following may also be used to assist engine startin extreme cold weather. After the last flight of
the day, drain the engine oil into a clean container so
the oil can be preheated. Cover the engine to prevent
ice or snow from collecting inside the cowling. When
preparing the aircraft for flight or engine runup after
these conditions have been followed, preheat the drainengine oil
ed
engineinstalled,
oil.
ceptacle
an external power source may be
connected to assist in cold weather or low battery
starting. Refer to paragraph 12-103 for use of the
external power receptacle.
Change 3

12-35

WARNING
Do not heat the oil above 121°C (250°F). A
flash fire may result. Before pulling the
propeller through, ascertain that the magneto switch is in the OFF position to prevent
accidental firing of the engine.
After preheating the engine oil, gasoline may be mixed with the heated oil in a ratio of 1 part gasoline to
12 parts engine oil before pouring into the engine oil
sump. If the free air temperature is below minus
29°C (-20F), the engine compartment should be preheated by a ground heater. After the engine compartment has been preheated, inspect all engine drain and
vent lines for presence of ice. After this procedure
has been complied with, pull propeller through several revolutions by hand before attempting to start
the engine.

CAUTION
Due to the desludging effect of the diluted
oil, engine operation should be observed
closely during the initial warm-up of the
engine. Engines that have considerable
amount of operational hours accumulated
since their last dilution period may be
seriously affected by the dilution process.
This will be caused by the diluted oil dilodging sludge and carbon deposits within
the engine. This residue will collect in the
oil sump and possibly clog the screened
inlet to the oil sump. Small deposits may
actually enter the oil sump and be trapped
by the main oil filter screen. Partial or
complete loss of engine lubrication may resuit from either condition. If these conditions are anticipated after oil dilution, the
engine should be run for several minutes
at normal operating temperatures and then
stopped and inspected for evidence of sludge
and carbon deposits in the oil sump and oil
filter screen. Future occurrence of this
condition can be prevented by diluting the
oil prior to each engine oil change. This
will also prevent the accumulation of the
sludge and carbon deposits.

SHOP NOTES:

12-36

Change 1

12-100. HOT WEATHER. Engine starting in hot
weather or with a hot engine is sometimes hampered
by vapor formation in the fuel lines. To purge the
vapor, move the mixture control to full rich, open
the throttle 1-1/2 inches and prime with the auxliary
fuel pump switch in the HI position until the fuel flow
indicator reads 4-6 gal/hr. Then shut off the fuel
pump switch and engage the starter. As the flooded
mixture becomes progressively leaner, reaching a
combustible mixture, the engine will start. If the
engine tends to die, turn the auxiliary fuel pump
switch momentarily to HI at appropriateintervals
until vapor is fully cleared and the engine runs
smoothly.
CAUTION
Never operate the starting motor more than
12 seconds at a time. Allow starter motor to

cool between cranking periods to avoid over-

heating. Longer cranking periods will shorten
the life of the starter motor.
12-101. SEACOAST AND HUMID AREAS. In salt
water areas special care should be taken to keep
the engine, accessories and airframe clean to prevent oxidation. In humid areas, fuel and oil should
be checked frequently and drained of condensation
to prevent corrosion.
12-102. DUSTY AREAS. Dust induced into the intake
system of the engine is probably the greatest single
cause of early engine wear. When operating in high
dust conditions, service the induction air filter daily
as outlined in Section 2. Also change engine oil and
lubricate airframe items more often than specified.
12-103. GROUND SERVICE RECEPTACLE. With
the ground service receptacle installed, the use of
an external power source is recommended for cold
weather starting, low battery starting and lengthy
maintenance of the aircraft electrical system. Refer
to Section 17 for additional information.
12-104. HAND-CRANKING. A normal hand-cranking procedure may be used to start the engine.

SECTION 12A
ENGINE
(TURBOCHARGED)

TABLE OF CONTENTS

*

Page

12A-2
ENGINE COWLING ............
12A-2
Description .............
12A-2
Removal and Installation .......
12A-2
Cleaning and Inspection ........
12A-2
Repair ...............
12A-2
.............
Cowl Flaps
. 12A-2
Description ........
12A-2
Removal and Installation .....
. 12A-2
...........
Rigging .
12A-2
................
ENGINE
12A-2
Description ........
12A-3
Engine Data ........
12A-3
..
..
(TBO)
Overhaul
Between
Time
12A-3
Overspeed Limitations ........
12A-4
Trouble Shooting ...........
. 12A-8
.............
Removal
12A-8A
Static Run-Up Procedures .......
. 12A-9
. ...
. ....
Cleaning
12A-9
.....
Accessories Removal .
. 12A-9
...
....
Inspection .
12A-9
........
..
Buld-Up
12A-9
.........
Installation
12A-10
Flexible Fluid Hoses ........
12A-10
Pressure Test ..........
. 12A-10
Replacement ........
12A-10
.....
....
Engine Baffles
12A- 10
Description ...........
12A-10
Cleaning and Inspection ......
. 12A-10
Removal and Installation ....
12A- 10
...........
.
Repair
12A-11
......
ENGINE OIL SYSTEM
. 12A-11
..
Description .......
12A-11
Trouble Shooting .........
12A-11
Full-Flow Oil Filter .......
12A-11
Description .........
Removal and Installation . . 12A-11
12A-11
......
.
Filter Adaptor .
12A-11
.......
Removal
Disassembly, Inspection and
12A-11
Reassembly ........
12A-11
Installation .........
12A-11
...........
Oil Cooler
. 12A-11
Description .......
12A-11
ENGINE FUEL SYSTEM ........
12A-11
........
Description .
. 12A-11
....
Fuel-Air Control Unit .
12A-11
Description .....
12A-11
.........
.
Removal

12A-14
Cleaning and Inspection ....
12A-14
.....
Installation ....
12A-14
.
Adjustments ........
12A-14
.
Fuel Manifold Valve.
12A-14
Description .........
12A-14
..........
Removal .
12A-14
....
Cleaning .
12A-14
.
...
Installation
12A-14
Fuel Discharge Nozzles ......
12A-14
Description ..
12A-14
.........
Removal .
. 12A- 14
Cleaning and Inspection .
12A-14
...
Installation ..
12A-14
Fuel Injection Pump ..
12A-14
Description ..
A-15
.
12
.....
Removal
12A-15
Installation .........
12A-15
Adjustment ........
Rigging Throttle Operated Micro12A..-15
Switch. ...
Auxiliary Electric Fuel Pump F\ov'
12A-15
Rate Adjustment .....
12A-16
INDUCTION AIR SYSTEM ..
.A-16
Description ....
12A-16
Airbox .......
.12A-16
Removal and Installation
12A- 16
Cleaning and Inspection ....
12A-16
..
Induction Air Filter
12A-16
Description .........
Removal and Installation . . . 12A-16
. 12A-16
Cleaning and Inspection .
12A-16
.........
IGNITION SYSTEM .
12A-16
Description .......
12A-16
Trouble Shooting .........
12A-16
.......
Magnetos .....
12A-16
..
.
Description ..
12A-16
..
Removal ....
12A-16
Internal Timing ......
Installation and Timing-to12A-16
Engine ..
12A-16
Maintenance ...
12A-16
.....
Magneto Check .
1ASpark Plugs ...........
12A-16
ENGINE CONTROLS .........
12A-16
..
Description ..
A-16
Rigging .............
12A-16
.
Throttle Control ..
12A-16
...
Mixture Control ..
. 12A-17
Propeller Control ...

Change 3

12A-1

12A-17
STARTING SYSTEM .........
12A- 17
Description ...........
12A-17
Trouble Shooting .......
.12A-17
Primary Maintenance ......
12A-17
.. .
Starter Motor ...
. . 12A-17
Removal and Installation
12A-17
EXHAUST SYSTEM ..........
12A-17
Description ...........
12A-17
............
Removal
12A-17
. ..
.. .
Installation ..
12A-20
.......
12A-21
TURBOCHARGER .........
12A-21
Description ...........
12A-21
Removal and Installation .....
CONTROLLER AND WASTE-GATE
. 12A-21
ACTUATOR ..........
12A-21
Functions ............

IInspection

12A-1.

12A-21
Operation ............
12o-24
Trouble Shooting .........
Controller and Turbocharger Operational
12A-26
.........
Flight Check
Removal and Installation of
12A-27
Turbocharger Controller ....
12A-27
......
Controller Adjustment
Removal and Installation of Waste12A-27
..
Gate and Actuator .
Adjustment of Waste-Gate Actuator. 12A-30
EXTREME WEATHER MAINTENANCE . 12A-30
12A-30
Cold Weather ..........
12A-30
Hot Weather ...........
12A-31
Seacoast and Humid Areas .....
12A-31
Dusty Areas ...........
12A-31
Ground Service Receptacle ......
12A-31
Hand-Cranking ............

12A-8.

ENGINE COWLING.

REMOVAL AND INSTALLATION.

Refer

to paragraph 12-8.
12A-2. DESCRIPTION. The engine cowling is similar to that described In Section 12, except it is wider
at the front, with additional ram air openings in the
right and left nose caps. The opening in the right
side supplies ram air to the turbocharger. The opening in the left side supplies ram air to the cabin heating system.
12A-3. REMOVAL AND INSTALLATION.
paragraph 12-3.

Refer to

12A-9. RIGGING. Refer to paragraph 12-9.
( Refer to figure 12-1)
12A-10.

ENGINE.

12A-11.

DESCRIPTION.

An air-cooled, horizon-

tally-opposed, direct-drive, fuel-Injected, six-cylinder turbocharged Continental TSIO-520 series engine,
driving a constant-speed propeller, is used to power

the aircraft. The cylinders, numbered from rear to
12A-4. CLEANING AND INSPECTION.
paragraph 12-4.

Refer to

Refer to paragraph 12-5.

12A-5.

REPAIR.

12A-6.

COWL FLAPS.

12A-7. DESCRIPTION. The cowl flaps are similar
to that described in Section 12, except the overboard
exhaust tube for the cabin heater extends through
the cutout in the aft portion of the left cowl flap.

SHOP NOTES:

12A-2

Change 1

front, are staggered to permit a separate throw on
the crankshaft for each connecting rod. The right
rear cylinder is number 1 and cylinders on the right
side are identified by odd numbers 1, 3 and 5. The
left rear cylinder is number 2 and the cylinders on
the left side are identified as 2, 4 and 6. Refer to
paragraph 12A-12 for engine data. For repair and
overhaul of the engine, accessories and propeller,
refer to the appropriate publications issued by their
manufacturer's. These publications are available
from the Cessna Service Parts Center.

12A-12.

ENGINE DATA.

Aircraft Series

TP206

TU206

Model (Continental)

TSIO-520-C

Same

BHP at RPM

285 at 2700

Same

Limiting Manifold Pressure
(Sea Level)

32.5 Inches Hg.

Same

Number of Cylinders

6-Horizontally Opposed

Same

Displacement
Bore
Stroke

520 Cubic Inches
5.25 Inches
5.00 Inches

Same
Same
Same

Compression Ratio

7.5:1

Same

Magnetos
Right Magneto

Slick Model No. 662
Fires 20° BTC Upper Right
and Lower Left
Fires 20° BTC Upper Left
and Lower Right

Same
Same

Firing Order

1-6-3-2-5-4

Same

Spark Plugs

18 MM (Refer to current Continental factory approved spark
plug chart.)
330±30 Lb-in.

Same

Fuel Metering System
Unmetered Fuel Pressure
~~~~ ~29

Continental Fuel Injection
6 to 7 PSI at 600 RPM
to 32 PSI at 2700 RPM

Same
Same
Same

Oil Sump Capacity
With Filter Element Change

12 U.S. Quarts
13 U.S. Quarts

Same
Same

Tachometer

Mechanical Drive

Same

Oil Pressure (PSI
Minimum Idling
Normal
Maximum (Cold Oil Starting)
Connection Location

10
30 to 60
100
Between No. 2 and No. 4 Cylinders

Same
Same
Same
Same

Oil Temperature
Normal Operating
Maximum Permissible
Probe Location

Within Green Arc
Red Line (240'F)
Below Oil Cooler

Same
Same
Same

Cylinder Head Temperature
Probe Location

Red Line (460°F) Max.
Lower Side No. 1 Cylinder

Same
Same

Approximate Dry Weight
With Accessories (Excluding
Turbocharger System)

483 Lb. (Weight is approximate
and will vary with optional
accessories installed.)

Same

Left Magneto

Torque

12A-12A. TIME BETWEEN OVERHAUL (TBO). Teledyne Continental Motors recommends engine overhaul
at 1400 hours operating time for the TSIO-520 series
engines. Refer to Continental Aircraft Engine Service
Bulletin M81-22 and to any superseding bulletins, revisions or supplements thereto. for further recom-

Same

Same

mendations. At the time of overhaul, engine accessories should be overhauled. Refer to Section 14 for
propeller and governor overhaul periods.
12A-12B OVERSPEED LIMITATIONS.
paragraph 12-12B.

Refer to

Change 2

12A-3

12A-13.

TROUBLE SHOOTING.

TROUBLE
ENGINE FAILS TO START.

PROBABLE CAUSE

REMEDY

Engine flooded or improper use
of starting procedure.

Use proper starting procedure.
Refer to Owner's Manual.

Defective aircraft fuel system.

Refer to Section 13.

Fuel tanks empty.

Service fuel tanks.

Spark plugs fouled or defective.

Remove, clean, inspect and regap.
Use new gaskets. Check cables
to presistently fouled plugs. Replace if defective.

Magneto impulse coupling failure.

Repair or install new coupling.

Defective magneto switch or
grounded magneto leads.

Repair or replace switch and leads.

Defective ignition system.

Refer to paragraph 12-79.

Induction air leakage.

Correct cause of air leakage.

Clogged fuel screen in fuel control
unit or defective unit.

Remove and clean.
defective unit.

Clogged fuel screen in fuel
manifold valve or defective

Remove and clean screen.
defective valve.

Replace
Replace

valve.

ENGINE STARTS BUT DIES, OR
WILL NOT IDLE PROPERLY.

12A-4

Clogged fuel injection lines or
discharge nozzles.

Remove and clean lines and nozzles.
Replace defective units.

Defective auxiliary fuel pump.

Refer to Section 13.

Engine-driven fuel pump not
permitting fuel from auxiliary
pump to bypass.

Install new engine-driven
fuel pump.

Vaporized fuel in system. (Most
likely to occur in hot weather with
a hot engine.)

Refer to paragraph 12A-114.

Propeller control in high pitch
(low rpm) position.

Use low pitch (high rpm) position
for all ground operations.

Improper idle speed or idle
mixture adjustment.

Refer to paragraph 12-46.

Defective aircraft fuel system.

Refer to Section 13.

Spark plugs fouled or defective.

Remove, clean, inspect and regap.
Use new gaskets. Check cables to
persistently fouled plugs. Replace
if defective.

Water in fuel system.

Drain fuel tank sumps, lines
and fuel strainer.

Defective ignition system.

Refer to paragraph 12-79.

12A-13.

TROUBLE SHOOTING (Cont).
TROUBLE

ENGINE STARTS BUT DIES, OR
WILL NOT IDLE PROPERLY
(CONT).

ENGINE HAS POOR ACCELERATION, RUNS ROUGHLY
AT SPEEDS ABOVE IDLE OR
LACKS POWER.

PROBABLE CAUSE

REMEDY

Induction air leakage.

Correct cause of air leakage.

Clogged fuel screen in fuel
control unit or defective unit.

Remove and clean.
defective unit.

Replace

Clogged fuel screen in fuel manifold valve or defective valve.

Remove and clean.
defective valve.

Replace

Restricted fuel injection lines
or discharge nozzles.

Remove, clean lines and nozzles.
Replace defective units.

Defective engine-driven fuel
pump.

Install and calibrate new pump.

Vaporized fuel in system.
(Most likely to occur in hot
weather with a hot engine.)

Refer to paragraph 12A-114.

Manual engine primer leaking.

Disconnect primer outlet line.
If fuel leaks through primer,
repair or replace primer.

Obstructed air intake.

Remove obstruction; service
air filter, if necessary.

Discharge nozzle air vent
manifolding restricted or
defective.

Check for bent lines or loose connections. Tighten loose connections. Remove restrictions and
replace defective components.

Defective engine.

Check compression and listen for
unusual engine noises. Check oil
filter for excessive metal. Repair
engine as required.

Idle mixture too lean.

Refer to paragraph 12-46.

Propeller control in high pitch
(low rpm) position.

Use low pitch (high rpm) position
for all ground operations.

Incorrect fuel-air mixture,
worn control linkage or
restricted air filter.

Replace worn elements of
control linkage. Service
air filter.

Defective ignition system.

Refer to paragraph 12-79.

Malfunctioning turbocharger.

Check operation, listen for unusual
noise. Check operation of wastegate valve and for exhaust system
defects. Tighten loose connections.

Improper fuel-air mixture.

Check intake manifold connections
for leaks. Tighten loose connections. Check fuel controls and linkage for setting and adjustment.

12A-5

12A-13.

TROUBLE SHOOTING (Cont).
TROUBLE

ENGINE HAS POOR ACCELERATION, RUNS ROUGHLY
AT SPEEDS ABOVE IDLE
OR LACKS POWER (CONT).

POOR IDLE CUT-OFF.

ENGINE LACKS POWER, REDUCTION IN MAXIMUM
MANIFOLD PRESSURE OR
CRITICAL ALTITUDE.

12A-6

PROBABLE CAUSE

REMEDY

Spark plugs fouled or defective.

Remove, clean, inspect and regap.
Use new gaskets. Check cables to
persistently fouled plugs. Replace
if defective.

Fuel pump pressure improperly
adjusted.

Refer to paragraph 12A-62.

Restriction in fuel injection
system.

Clean out restriction.
defective items.

Propeller out of balance.

Check and balance propeller.

Defective engine.

Check compression, check oil
filter for excessive metal.
Listen for unusual noises.
Repair engine as required.

Exhaust system leakage.

Refer to paragraph 12A-99.

Turbocharger wheels rubbing.

Replace turbocharger.

Improperly adjusted or defective
waste-gate controller.

Refer to paragraph 12A-111.

Leak in turbocharger discharge
pressure system.

Correct cause of leaks. Repair
or replace damaged parts.

MrnifolA pressure overshoot.
(Most likely to occur when
engine is accelerated too
rapidly.)

Move throttle about two-thirds
open. Let engine accelerate
and peak. Move throttle to
full open.

Engine oil viscosity too high
for ambient air.

Refer to Section 2 for proper
grade of oil.

Mixture control linkage improperly rigged.

Refer to paragraph 12-86.

Defective or dirty fuel manifold
valve.

Remove and clean manifold
valve.

Fuel contamination.

Drain all fuel and flush out fuel
system. Clean all screens, fuel
strainers, fuel manifold valves,
nozzles and fuel lines.

Defective mixture control
valve in fuel pump.

Replace fuel pump.

Incorrectly adjusted throttle
control, "sticky" linkage or
dirty air filter.

Check movement of linkage by moving control through range of travel.
Make proper adjustments and replace worn components. Service
air filter.

Replace

12A-13.

TROUBLE SHOOTING (Cont).
TROUBLE

ENGINE LACKS POWER, REDUCTION IN MAXIMUM
MANIFOLD PRESSURE OR
CRITICAL ALTITUDE (CONT).

PROBABLE CAUSE

REMEDY

Defective ignition system.

Inspect spark plugs for fouled
electrodes, heavy carbon deposits, erosion of electrodes,
improperly adjusted electrode
gaps and cracked porcelains.
Test plugs for regular firing
under pressure. Replace damaged or misfiring plugs.

Improperly adjusted waste-gate
valve.

Refer to paragraph 12A-111.

Loose or damaged exhaust
system.

Inspect entire exhaust system to
turbocharger for cracks and
leaking connections. Tighten
connections and replace damaged
parts.

Loose or damaged manifolding.

Inspect entire manifolding system
for possible leakage at connections.
Replace damaged components,
tighten all connections and clamps.

Fuel discharge nozzle defective.

Inspect fuel discharge nozzle vent
manifolding for leaking connections.
Tighten and repair as required.
Check for restricted nozzles and
lines and clean and replace as
necessary.

Malfunctioning turbocharger.

Check for unusual noise in turbocharger. If malfunction is suspected, remove exhaust and/or
air inlet connections and check rotor assembly, for possible rubbing
in housing, damaged rotor blades
or defective bearings. Replace
turbocharger if damage is noted.

BLACK SMOKE EXHAUST.

Turbo coking, oil forced through
seal of turbine housing.

Clean or change turbocharger.

HIGH CYLINDER HEAD
TEMPERATURE.

Defective cylinder head temperature indicating system.

Refer to Section 16.

Improper use of cowl flaps.

Refer to Owner's Manual.

Engine baffles loose, bent or
missing.

Install baffles properly.
replace if defective.

Dirt accumulated on cylinder
cooling fins.

Clean thoroughly.

Incorrect grade of fuel.

Drain and refill with proper fuel.

Repair or

12A-7

12A-13.

TROUBLE SHOOTING (Cont).

TROUBLE

PROBABLE CAUSE

HIGH CYLINDER HEAD
TEMPERATURE (CONT).

REMEDY

Incorrect ignition timing.

Refer to paragraph 12-78.

Improper use of mixture control.

Refer to Owner's Manual.

Defective engine.

Repair as required.

HIGH OR LOW OIL
TEMPERATURE
OR PRESSURE.

Refer to paragraph 12-30.

NOTE
Refer to paragraph 12A-106 for trouble shooting of controller
and waste-gate actuator.

12A-14. REMOVAL. If an engine is to be placed in
storage or returned to the manufacturer for overhaul, proper preparatory steps should be taken for
corrosion prevention prior to beginning the removal
procedure. Refer to Section 2 for storage preparation. The following engine removal procedure is
based upon the engine being removed from the aircraft as a complete unit with the turbocharger and
accessories installed.
NOTE
Tag each item when disconnected to aid in
identifying wires, hoses, lines and control
linkages when engine is reinstalled. Likewise, shop notes made during removal will
often clarify reinstallation. Protect openings, exposed as a result of removing or
disconnecting units, against entry of foreign
material by installing covers or sealing with
tape.
a. Place all cabin switches in the OFF position.
b. Place fuel selector valve in the OFF position.
c. Remove engine cowling in accordance with paragraph 12-3.
d. Disconnect battery cables and insulate terminals
as a safety precaution. Remove battery and battery
box for additional clearance, if desired.
e. Drain fuel strainer and lines with strainer drain
control.
NOTE
During the following procedures, remove
any clamps or lacings which secure controls, wires, hoses or lines to the engine,
engine nacelle or attached brackets, so
they will not interfere with engine removal.
Some of the items listed can be disconnected
at more than one place. It may be desirable
12A-8

to disconnect some of these items at other
than the places indicated. The reason for
engine removal should be the governing factor in deciding at which point to disconnect
them. Omit any of the items which are not
present on a particular engine installation.
f. Drain the engine oil sump and oil cooler.
g. Disconnect magneto primary lead wires at
magnetos.

IWARNING
The magnetos are in a SWITCH ON condition
when the switch wires are disconnected.
Ground the magneto points or remove the high
tension wires from the magnetos or spark
plugs to prevent accidental firing.
h. Remove the spinner and propeller in accordance
with Section 14. Cover exposed end of crankshaft
flange and propeller flange to prevent entry of foreign
material.
i. Disconnect throttle, mixture and propeller controls from their respective units. Remove clamps
attaching controls to engine and pull controls aft
clear of engine. Use care to avoid bending controls
too sharply. Note EXACT position, size and number
of attaching washers and spacers for reference on
reinstallation.
j. Disconnect wires and cables as follows:
1. Disconnect tachometer drive shaft at adapter.

CAUTION
When disconnecting starter cable do not
permit starter terminal bolt to rotate.
Rotation of the bolt could break the conductor between bolt and field coils causing the starter to be inoperative.

2. Disconnect starter electrical cable at starter.
3. Disconnect cylinder head temperature wire at
probe.
4. Disconnect oil temperature wire at probe below oil cooler.
5. Disconnect electrical wires and wire shielding ground at alternator.
6. Disconnect exhaust gas temperature wires at
quick-disconnects.
7. Disconnect electrical wires at throttle microswitch.
8. Remove all clamps and lacings attaching
wires or cables to engine and pull wires and cables
aft to clear engine.
k. Disconnect lines and hoses as follows:
1. Disconnect vacuum hose at vacuum pump
and remove oil separator vent line.
NI

aircraft checking for any items which would interfere
with the engine removal. Balance the engine by hand
and carefully guide the disconnected parts out as the
engine is removed.
p. Remove engine shock-mounts

WARNING

Residual fuel and oil draining from disconnected lines and hoses constitutes a fire
hazard. Use caution to prevent accumulation of such fuel and oil when lines or hoses
are disconnected.
2. Disconnect fuel supply and vapor return
hoses at fuel pump. Disconnect and remove fuel
pump drain line.
3. Disconnect manifold pressure line at
intake manifold.
4. Disconnect the fuel-flow gage line at firewall.
5. Disconnect the oil pressure line at the
engine.
6. Disconnect and remove the right and left
manifold drain lines and the balance tube drain line.
7. Disconnect air and oil lines at the waste-gate
controller, located on the firewall.
8. Disconnect the air vent line to fuel-flow gage.
at firewall.
9. Disconnect engine primer lines at right and
left intake manifolds.
10. Disconnect the oil drain line from oil deflector under external oil filter.
I. Carefully check the engine again to ensure ALL
hoses, lines, wires, cables, clamps and lacings are
disconnected or removed which would interfere with
the engine removal. Ensure all wires, cables and
engine controls have been pulled aft to clear the engine.

|CT~AC~UTION
I
~equipped

CAUTION

I

Place a suitable stand under tail tie-down
ring before removing engine. The loss of
engine weight will cause the aircraft to be
tail heavy.
m. Attach a hoist to the lifting lug at the top center
of the engine crankcase. Lift engine just enough to
relieve the weight from the engine mounts.
n. Remove mount bolts, ground strap and heat
shields.
o. Slowly hoist engine out of nacelle and clear of

NOTE
If shock-mounts will be re-used, mark each
one so it will be reinstalled in exactly the
same position. If new shock-mounts will be
installed, position them as illustrated in figure
12-2.
12A-14A. STATIC RUN-UP PROCEDURES. In a
case of suspected low engine power, a static RPM
run-up should be conducted as follows:
a
. Run-up engine, using take-off power and mixtlure settings, with the aircraft facing 90 ° right and
then left to the wind direction.
b. Record the RPM obtained in each run-up position.
NOTE
Daily changes in atmospheric pressure, temture and humidity will have a slight effect
on static rup-up.
c. Average the results of the RPM obtained. It
should be within 50 RPM of 2650 RPM.
d. If the average results of the RPM obtained are
lower than stated above, the following recommended
checks may be performed to determine a possible
deficiency.
1. Check governor control for proper rigging.
It should be determined that the governor control
arm travels to the high RPM stop on the governor
and that the high RPM stop screw is adjusted properly. ( Refer to Section 14 for procedures).
NOTE
If verification of governor operation is necessary the governor may be removed from the
engine and a flat plate installed over the engine pad. Run-up engine to determine that
governor was adjusted properly.
2.

Check carburetor heat control (carburetor
engines) for proper rigging. If partially
open it would cause a slight power loss. On fuel injected engines check operation of alternate air door
spring or magnetic lock to make sure door will remain closed in normal operation.
3. Check magneto timing, spark plugs and ignition harness for settings and conditions.
4. On fuel injection engines, check fuel injection
nozzles for restriction and check for correct unmetered fuel flow.
5. Check condition of induction air filter. Clean
if required.
6. Perform an engine compression check (Refer
to engine Manufacturer's Manual).

Change 1

12A-8A/(12A-8B blank)

12A-15.

CLEANING.

Refer to paragraph 12-15.

12A-16. ACCESSORIES REMOVAL.
graph 12-16.
12A-17.

INSPECTION.

12A-18.

BUILD-UP.

Refer to para-

Refer to paragraph 12-17.

NOTE
Be sure engine shock-mounts, spacers
and washers are in place as the engine is
lowered into position.

Refer to paragraph 12-18.

12A-19. INSTALLATION. Before installing the
engine on the aircraft, install any items which were
removed from the engine or aircraft after the engine
was removed.
NOTE
Remove all protective covers,
and identification tags as each
nected or installed. Omit any
present on a particular engine

plugs, caps
item is conitems not
installation.

a. Hoist the engine to a point just above the nacelle.
b. Install engine shock-mounts and ground strap as
illustrated in figure 12-2.
c. Carefully lower engine slowly into place on the
engine mounts. Route controls, lines, hoses and
wires in place as the engine is positioned on the engine mounts.

d. Attach ground strap under engine sump bolt
and install engine mount bolts. Torque bolts to 300+
50-00 lb-in. Bend tab washers to form lock for mount
bolts. Install heat shields.
e. Remove support stand placed under tail tie-down
fitting and remove hoist.
NOTE
If the exhaust system was loosened or removed,
refer to paragraph 12A-98.

f. Connect flexible ducting on heater shroud and
cabin valve.
g. Route propeller governor control along left side
of engine and secure with clamps.

SHOP NOTES:

Change 1

12A-9

NOTE
Throughout the aircraft fuel system, from
the fuel cells to the engine driven fuel pump,
use RAS-4 (Snap-On Tools Corp., Kenosha,
Wisconsin), MIL-T-5544 (Thread Compound,
Antiseize, Graphite-Petrolatum) or equivalent,
as a thread lubricant or to seal a leaking connection. Apply sparingly to male fittings only, omitting the first two threads. Always ensure that a

k. Rig engine controls in accordance with paragraphs 12-85, 12-86 and 12-87.
1. Install propeller and spinner in accordance with
instructions outlined in Section 14.
m. Complete a magneto switch ground-out and continuity check, then connect primary lead wires to the
magnetos. Remove the temporary ground or connect
spark plug leads, whichever procedure was used during removal.
WAR
I

compound, the residue from a previously used
compound or any other foreign material cannot
enter the system. Throughout the fuel injection system, from the engine-driven fuel pump
through the discharge nozzles, use only I fuel
soluble lubricant, such as engine lubricating
oil, on fitting threads. Do not use any other
form of thread compound on the fuel injection
system fittings.
h.

Connect lines and hoses as follows:
1. Install and connect the left and right manifold
drain lines and the balance tube drain line.
2. Connect the oil pressure line at its fitting.
3. Connect the fuel-flow gage line at firewall.
4. Connect the fuel supply and the vapor return
lines at the fuel pump. Connect and install fuel pump
drain line.
5. Connect manifold pressure line at intake manifold.
6. Connect vacuum line at the vacuum pump, and
Install oil separator vent line.
7. Connect air and oil lines at waste-gate controller on firewall.
8. Connect air vent line to fuel-flow gage line
at firewall.
9. Connect engine primer lines at right and left
intake manifolds.
10. Connect oil drain line to oil deflector under
external oil filter.
11. Install all clamps securing lines and hoses to
engine or structure.
i. Connect wires and cables as follows:
1. Connect oil temperature wire at probe below
oil cooler.
2. Connect tachometer drive to adapter and torque to 100 lb-in.

IWARNING

WARNING|

]

Be sure magneto switch is in OFF position
when connecting switch wires to magnetos.
n. Clean and install induction air filter in accordance with Section 2.
o. Service engine with proper grade and quantity of
engine oil. Refer to Section 2 if engine is new, newly
overhauled or has been in storage.
p. Check all switches are in the OFF position and
connect battery cables.
q. Inspect engine installation for security, correct
routing of controls, lines, hoses and electrical wiring, proper safetying and tightness of all components.
NOTE
When installing a new or newly overhauled
engine, and prior to starting the engine,
disconnect the oil inlet line at the controller
and the oil outlet line at the controller.
Connect these oil lines to a full-flow oil
filter, allowing oil to bypass the controller.
With filter connected, operate engine approximately 15 minutes to filter out any foreign
particles from the oil. This is done to prevent foreign material from entering the controller.
r. Install engine cowling in accordance with paragraph 12-3.
s. Perform an engine run-up and make final adjustments on the engine controls.
12A-20. FLEXIBLE FLUID HOSES.
graph 12-20.

12A-21.

PRESSURE TEST.

Refer to para-

Refer to paragraph 12-

21.
When connecting starter cable, do not permit
starter terminal bolt to rotate. Rotation of
the bolt could break conductor between terminal
and field coils causing starter to be inoperative.
3. Connect starter electrical lead.
4. Connect cylinder head temperature wire at
probe.
5. Connect electrical wires and wire shielding
ground to alternator.
6. Connect electrical wires to throttle switch.
7. Connect exhaust gas temperature wires at
quick-disconnects.
8. Install clamps that attach wires or cables, to
engine or structure.
j. Connect engine controls and install block clamps.
12A-10

12A-22.
22.

REPLACEMENT.

Refer to paragraph 12-

12A-23.
12-23.

ENGINE BAFFLES.

12A-24.

DESCRIPTION.

Refer to paragraph

Refer to paragraph 12-24.

12A-25. CLEANING AND INSPECTION.
paragraph 12-25.
12A-26. REMOVAL AND INSTALLATION.
paragraph 12-26.
12A-27.

REPAIR.

Refer to

Refer to

Refer to paragraph 12-27.

12A-28.
12A-1.

ENGINE OIL SYSTEM.

Refer to figure

12A-29. DESCRIPTION. The engine lubrication
system is a full-pressure, wet-sump type. Lubricating oil is drawn from the engine sump to the oil
pump through a suction screen and tube. From the
pump, oil under pressure is passed to the full-flow
oil filter, where it is filtered before entering the
passages of the engine. Bypass valves are provided.
Oil from the filter is routed through drilled and cored passages to all moving parts requiring lubrication.
Oil furnished to the propeller governor for propeller
operation is also routed through internal passages.
Oil pressure is maintained by an adjustable, springloaded relief valve mounted in the lower portion of
the pump body. Oil temperature is automatically
regulated by an oil cooler and a thermostat control
valve. When +he oil temperature reaches a predetermined temperature the thermostat valve closes,
causing the oil to be routed through the externally
mounted cooler. Engine oil is also used to control
the waste-gate and lubricate the turbocharger bearings. Oil is returned to the engine sump from the
turbocharger by a scavenger pump, which is integral with the engine oil pump. The oil filler neck is
located on top of the engine and is reached through
an access door in the top of the left cowl. The oil
level in the sump is checked on a dipstick at the
rear of number two cylinder and is reached through
an access door in the side of the left cowl.
12A-30.
12-30.

TROUBLE SHOOTING.

Refer to paragraph

12A-31. FULL-FLOW OIL FILTER.
graph 12-31.
12A-32.

DESCRIPTION.

Refer to para-

Refer to paragraph 12-32.

12A-33. REMOVAL AND INSTALLATION.
to paragraph 12-33.
12A-34.
12-34.

FILTER ADAPTER.

12A-35.

REMOVAL.

Refer to paragraph 12-35.

12A-36. DISASSEMBLY, INSPECTION AND REASSEMBLY. Refer to paragraph 12-36.
INSTALLATION.

12A-38.

OIL COOLER.

12A-39.

DESCRIPTION.

12A-40.
12A-2.

ENGINE FUEL SYSTEM.

NOTE
Throughout the aircraft fuel system, from
the fuel cells to the engine-driven fuel pump,
use RAS-4 (Snap-On Tools Corp., Kenosha,
Wisconsin), MIL-T-5544 (Thread Compound,
Antiseize, Graphite-Petrolatum) or equivalent,
as a thread lubricant or to seal a leaking connection. Apply sparingly to male fittings only,
omitting the first two threads. Always ensure
that a compound, the residue from a previously
used compound or any other foreign material
cannot enter the system. Throughout the fuel
injection system, from the engine-driven fuel
pump through the discharge nozzles, use only
a fuel soluble lubricant, such as engine lubricating oil, on the fitting threads. Do not use
any other form of thread compound on the injection system fittings.
12A-42. FUEL-AIR CONTROL UNIT.
paragraph 12-42.

Refer to

Refer

Refer to paragraph

12A-37.

engine airflow. A manual mixture control and a fuel
flow indicator are provided for leaning at any combination of altitude and power setting. The fuel flow
indicator is calibrated in gallons per hour and indicates approximately the gallons of fuel consumed per
hour. The continuous-flow system uses a typical
rotary vane fuel pump. There are no running parts
in this system except for the engine-driven fuel pump.
The four major components of the system are: the
fuel injection pump, fuel-air control unit, fuel manifold valve and the fuel discharge nozzles. The fuel
injection pump incorporates an adjustable aneroid
sensing unit which is pressurized from the discharge
side of the turbocharger compressor. Turbocharger
discharge air pressure is also used to vent the fuel
discharge nozzles and the vent port of the fuel-flow
gage.

Refer to paragraph 12-37.
Refer to paragraph 12-38.
Refer to paragraph 12-39.
Refer to figure

12A-41. DESCRIPTION. The fuel injection system
is a low pressure system of injecting fuel into the
intake valve port of each cylinder. It is a multinozzle, continuous-flow type which controls fuel
flow to match engine airflow. Any change in throttle
position, engine speed, or a combination of both,
causes changes in fuel flow in the correct relation to

12A-43.

DESCRIPTION.

Refer to paragraph 12-43.

12A-44. REMOVAL.
a. Place all cabin switches and fuel shut-off valve
in the OFF position.
b. Remove cowling in accordance with paragraph
12-3.
c. Loosen clamp and disconnect flexible duct from
elbow at top of air throttle.
d. Tag and disconnect electrical wires from electric fuel pump microswitch.
e. Disconnect throttle and mixture control rod ends
at fuel-air control unit.
NOTE
Cap or plug all disconnected hoses, lines and
fittings.
f. Disconnect cooling air blast tube from fuel control valve shroud.
g. Disconnect and tag all fuel lines at the fuel control valve.
h. Remove nuts and washers securing triangular
brace to fuel-air control unit and engine, at lower
end of control unit. Remove brace.
12A-11

CODE:

PLUG

THERMOSTAT

PRESSURE OIL
SUMP OIL, RETURN
OIL AND SUCTION
OIL

TO
PROPELLER
PLUG

PROPELLER
CONTROL

PROPELLER
GOVERNOR

THERMOSTAT
CLOSED

OIL
TEMPERATURE
GAGE

OIL
FILLER
CAP
OIL
PRESSURE
GAGE

OIL PRESSURE

OIL DIPSTICK
OIL SUMP
DRAIN PLUG
FUEL LINE FROM
OPTIONAL OIL
DILUTION SYSTEM

RELIEF VALVE
OIL PUMP
BYPASS
VALVE

ENGINE
OIL
PUMP

OIL FILTER
BYPASS VALVE
FILTER

CHECK VALVE
SCAVENGER

OUTLET

PUMP
TURBOCHARGER

WASTE-GATE
CONTROLLER

CHECK VALVE
EXTERNAL
OIL FILTER

Figure 12A-1.
12A -12

Oil System Schematic

WASTE-GATE
ACTUATOR

Manifold Valve
To Fuel Flow Gage

--------

To Vent Port of
Fuel Flow Gage

Air From
-

-Turbocharger
1
\Discharge

fl

Fuel Inlet
LEGEND:
RELIEF
VALVE PRESSURE
RETURN FUEL
METERED FUEL
Orifice

PUMP PRESSURE
INLET PRESSURE
TURBOCHARGER AIR DISCHARGE PRESSURE

Figure 12A-2.

Injection Mixture Outlet

Fuel System Schematic
12A-13

i. Remove bolt attaching fuel-air control unit to
brace at top of control unit.
j. Loosen hose clamps which secure fuel-air control unit to right and left intake manifold assemblies
and slip hoses from fuel-air control unit.
k. Remove fuel-air control unit.
12A-45. CLEANING AND INSPECTION.
paragraph 12-45.

Refer to

12A-46. INSTALLATION.
a. Place control unit in position at rear of engine.
b. Install bolt attaching control unit to brace at top
of unit. Ascertain that shock-mount is in place and in
good condition.
c. Install triangular brace at lower end of control
unit.
d. Install hoses and clamps which secure control
unit to right and left intake manifold assemblies.
Tighten hose clamps.
e. Connect fuel lines to unit and connect air blast
tube at fuel control shroud.
f. Connect throttle and mixture control rod ends
to control unit.
g. Connect electrical wires to electric fuel pump
microswitch. Check switch rigging in accordance
with Section 13.
h. Install induction air duct to elbow at top of control unit.
i. Inspect installation and install cowling.
12A-47. ADJUSTMENTS.

Refer to paragraph 12-46.

12A-48. FUEL MANIFOLD VALVE (FUEL DISTRIBUTOR). Refer to paragraph 12-47.
12A-49.

DESCRIPTION.

12A-50.

REMOVAL.

Refer to paragraph 12-49.

12A-51.

CLEANING.

Refer to paragraph 12-50.

12A-52.

INSTALLATION.

12A-53.

FUEL DISCHARGE NOZZLES.

Refer to paragraph 12-48.

Refer to paragraph 12-51.

12A-54. DESCRIPTION. From the fuel manifold
valve, individual, identical size and length fuel lines
carry metered fuel to the fuel discharge nozzles located in the cylinder heads. The outlet of each nozzle
is directed into the intake port of each cylinder. An
air bleed and nozzle pressurization arrangement is
incorporated in each nozzle to aid in vaporization of
the fuel. The nozzles are calibrated in several ranges.
All nozzles furnished for one engine are of the same
calibrated range and are identified by a number and
suffix letter stamped on the flat portion of the nozzle
body. When replacing a fuel discharge nozzle, be
sure that it is of the same calibrated range as the
rest of the nozzles in that engine. When a complete
set of nozzles is being replaced, the number must be
the same as the one removed but the suffix letter
may be different, as long as they are the same for
all nozzles being installed in a particular engine.
12A-55. REMOVAL.
a. Remove engine cowling in accordance with para12A-14

Change 3

graph 12-3.
NOTE
Plug or cap all disconnected lines and fittings.
b. Disconnect nozzle pressurization line at nozzles
and disconnect pressurization line at "tee" fitting so
that pressurization line may be moved away from
discharge nozzles.
c. Disconnect fuel injection line at fuel discharge
nozzle.
d. Using care to prevent damage or loss of washers
and O-rings, lift sleeve assembly from fuel discharge nozzle.
e. Using a standard 1/2-inch deep socket, remove
fuel discharge nozzle from cylinder.
12A-56. CLEANING AND INSPECTION.
paragraph 12-55.

Refer to

12A-57. INSTALLATION.
a. Using a standard 1/2-inch deep socket, install
nozzle body in cylinder and tighten to a torque value
of 60-80 lb-in.
b. Install O-rings, sleeve assembly and washers.
c. Align sleeve assembly and connect pressurization line to nozzles. Connect pressurization line to
"tee" fitting.
d. Install O-ring and washer at top of discharge
nozzle and connect fuel injection line to nozzle.
e. Inspect installation for crimped lines and loose
fittings.
f. Inspect nozzle pressurization vent system for
leakage. A tight system is required, since turbocharger discharge pressure is applied to various
other components of the injection system.
g.

Install cowling.

12A-58. FUEL INJECTION PUMP.
12A-59. DESCRIPTION. The fuel pump is a positive
displacement, rotating vane type. It has a splined
shaft for connection to the accessory drive section
of the engine. Fuel enters the pump at the swirl well
of the pump vapor separator. Here, vapor is separated by a swirling motion so that only liquid fuel is fed
to the pump. The vapor is drawn from the top center
of the swirl well by a small pressure jet of fuel and
is fed into the vapor return line where it is returned
to the fuel tank. Since the pump is engine-driven,
changes in engine speed affect total pump flow proportionally. A check valve allows the auxiliary fuel pump
pressure to bypass the engine-driven pump for starting, or in the event of engine-driven fuel pump failure
in flight. The pump supplies more fuel than is required
by the engine; therefore, a relief valve is provided
to maintain a constant fuel pump pressure. The
engine-driven fuel pump is equipped with an aneroid.
The aneroid and relief valve are pressurized from the
discharge side of the turbocharger compressor to
maintain a proper fuel/air ratio at altitude. The
aneroid is adjustable for fuel pump outlet pressure
at full throttle and the relief valve is adjustable for
fuel pump outlet pressure at idle.

12A-60. REMOVAL.
a. Place fuel selector valve handle in OFF position.
b. Remove engine cowling in accordance with paragraph 12-3.
c. Remove alternator and left rear intake elbow.
d. Hoist engine far enough to remove weight from
engine mount and remove left rear engine mount leg,
shock-mount and alternator bracket.
e. Remove flexible duct and shroud, removing fuel
lines and fittings as necessary. Tag each fitting and
line for identification and cap or seal to prevent entry of foreign material. Flanges of shroud may be
straightened to facilitate removal and installation,
but must be re-formed after installation. Note angular position of fittings before removal.
f. Remove nuts and washers attaching fuel pump
to engine and pull pump aft to remove. Remove thin
gasket.
g. Place temporary cover on pump mounting pad.
12A-61. INSTALLATION.
a. Install and align any fittings removed after pump
removal.
b. Using new thin gasket, install pump with aneroid
chamber down.
c. Install cooling shroud and remainder of fittings,
bending flanges of shroud to their original positions
and aligning fittings as noted during removal.
d. Connect all fuel lines and shroud flexible duct.
e. Install alternator bracket, shock-mount and
engine mount leg. Remove hoist, then adjust alternator drive belt tension. Refer to Section 17.
f. Install intake elbow.
g. Start engine and perform an operational check,
adjusting fuel pump if required.
h. Install cowling.
12A-62. ADJUSTMENT. Adjustments of the fuel injection pump requires special equipment and procedures. Adjustment to the aneroid applies only to the
full throttle setting. Adjustment of the idle position
is obtained through the relief valve. To adjust the
pump to the pressures specified in paragraph 12A-12,
proceed as follows:
a. Remove engine cowling in accordance with paragraph 12-3.
b. Disconnect the existing engine-driven fuel pump
pressure hose at the fuel metering unit and the existing fuel gage vent hose at the air manifold valve.
Connect the test gage pressure hoses. vent hose and
fittings into the fuel injection system as illustrated in
figure 12A-8.
c. The test gage MUST be vented to upper deck
pressure and MUST be held as near to the level of
the engine-driven pump as possible. Bleed air from
test gage line prior to taking readings.
NOTE
Cessna Service Kit No. SK320-2 provides
a test gage, lines and fittings for connecting the test gage into the system to
perform accurate calibration of the enginedriven fuel pump.

NOTE
The test gage should be checked for accuracy
at least every 90 days or anytime an error is
suspected. The tachometer accuracy should
also be determined prior to making any adjustments to the pump.
d. Start engine and warm-up thoroughly. Set mixture control to full rich position and propeller control full forward (low pitch, high rpm).
e. Adjust engine idle speed to 600 ± 25 rpm and
check test gage for 6-7 PSI. Refer to figure 12-7
for idle mixture adjustment.
NOTE
Do not adjust idle mixture until idle pump
pressure is obtained.

WARNING
DO NOT make fuel pump pressure adjustments while engine is operating.
f. If the pump pressure is not 6 to 7 PSI, stop engine and turn the fuel pump relief valve adjustment,
on the centerline of the fuel pump clockwise (CW) to
increase pressure and counterclockwise (CCW) to
decrease pressure.
g. Maintaining idle pump pressure and idle RPM,
obtain correct idle mixture in accordance with paragraph 12-46.
h. Completion of the preceding steps have provided:
I. Correct idle pump pressure.
2. Correct fuel flow.
3. Correct fuel metering cam to throttle plate
orientation.
i. Advance to full throttle and maximum rated engine speed with the mixture control in full rich position and propeller control in full forward (low pitch,
high rpm).
j. Check test gage for pressures specified in paragraph 12A-12. If pressure is incorrect, stop engine
and adjust pressure by loosening locknut and turning
the adjusting screw located at rear of aneroid counterclockwise (CCW) to increase pressure and clockwise (CW) to decrease pressure.
NOTE
If at static run-up, rated RPM cannot be
achieved at full throttle, adjust pump pressure slightly below limits making certain
the correct pressures are obtained when
rated RPM is achieved during take-off roll.
k. After correct pressures are obtained, tighten
locknut.
1. Remove test equipment, run engine to check for
leaks and install cowling.
12A-62A. RIGGING THROTTLE OPERATED MICROSWITCH. Refer to Section 13,
12A-62B. AUXILIARY ELECTRIC FUEL PUMP
FLOW RATE ADJUSTMENT. Refer to Section 13.

D2007C3-13 Temporary Change 2
22 February 1978

Change 3

12A-15

12A-63.

INDUCTION AIR SYSTEM.

12A-64. DESCRIPTION. Ram air to the engine enters an induction air duct at the right side of the nose
cap. The air is filtered through a dry filter, located
in the induction airbox. From the filter, the air passes through a flexible duct to the inlet of the turbocharger compressor. The pressurized air is then
routed through a duct to the fuel-air control unit
mounted behind the engine and is then supplied to
the cylinders through the intake manifold piping. The
fuel-air control unit is connected to the cylinder intake manifold by elbows, hoses and clamps. The intake manifold is attached to each cylinder by four
bolts through a welded flange, which is sealed by a
gasket. A balance tube passes around the front side
of the engine to complete the manifold assembly. An
alternate air door, mounted in the duct between the
filter and the turbocharger compressor, is held closed by a small magnet. If the induction air filter
should become clogged, suction from the turbocharger
compressor will open the door permitting the compressor to draw heated, unfiltered air from within
the engine compartment. The alternate air door
should be checked periodically for freedom of operation and complete closing. The induction air filter
should be removed and cleaned at each 50-hour inspection, more often when operating under dusty conditions. Refer to Section 2.
12A-65.

AIRBOX.

12A-66. REMOVAL AND INSTALLATION.
a. Remove engine cowling in accordance with paragraph 12-3.
b. Loosen clamp at lower end of airbox and remove
flexible duct.
c. Remove two screws, washers and nuts attaching
airbox to upper rear engine baffle.
d. Remove four screws attaching airbox to induction air duct and work airbox and filter from duct.
e. Remove screws attaching clips on duct to clips
on rocker box covers.
f. Remove screws attaching lower side of induction
air duct to the two front cylinder rocker box covers.
g. Loosen clamp and remove air duct from flexible
inlet air duct and remove duct.
h. Reverse the preceding steps for reinstallation.

b. Remove screws attaching airbox to upper rear
baffle.
c. Loosen clamp and disconnect flexible air duct to
airbox.
d. Remove four screws attaching airbox to forward
air duct and work airbox and filter from aircraft.
e. Remove four bolts, washers and nuts attaching
filter between airbox halves.
NOTE
When installing filter, note direction of air
flow. Inspect and install gasket at aft face
of filter assembly. Also, when tightening
bolts fastening filter, push inward on lower
end of the upper duct (where turbocharger
inlet connects to the upper duct). This is
done so that inlet hose doesn't chafe against
the cowling.
f.

Reverse the preceding steps for reinstallation.

12A-71. CLEANING AND INSPECTION. Clean and
inspect filter in accordance with Section 2.
12A-72.
12-71.

IGNITION SYSTEM.

12A-73.

DESCRIPTION.

12A-74.
12-73.

TROUBLE SHOOTING.

12A-75.

MAGNETOS.

12A-76.

DESCRIPTION.

12A-77.

REMOVAL.

12A-78.
12-77.

INTERNAL TIMING.

Refer to paragraph

Refer to paragraph 12-72.
Refer to paragraph

Refer to paragraph 12-74.
Refer to paragraph 12-75.

Refer to paragraph 12-76.
Refer to paragraph

12A-79. INSTALLATION AND TIMING-TO-ENGINE.
Refer to paragraph 12-78.
12A-80.

MAINTENANCE.

12A-81.
12-80.

MAGNETO CHECK.

12A-82.

SPARK PLUGS.

12A-83.
12-82.

ENGINE CONTROLS.

12A-67. CLEANING AND INSPECTION. Refer to
paragraph 12-66.

12A-84.

DESCRIPTION.

12A-68.

12A-85.

RIGGING.

12A-86.
12-85.

THROTTLE CONTROL.

12A-87.
12-86.

MIXTURE CONTROL.

Refer to paragraph 12-79.
Refer to paragraph

NOTE
Clean filter and ascertain that induction air
ducts and airbox are clean when installing.

INDUCTION AIR FILTER.

12A-69. DESCRIPTION. An induction air filter,
mounted in the aft end of the airbox removes dust
particles from the ram air entering the engine.
12A-70. REMOVAL AND INSTALLATION.
a. Remove right half of engine cowling in accordance with paragraph 12-3.
12A-16

Refer to paragraph 12-81.
Refer to paragraph

Refer to paragraph 12-83.

Refer to paragraph 12-84.
Refer to paragraph
Refer to paragraph

12A-88.
14.

PROPELLER CONTROL.

12A-89.
12-88.

STARTING SYSTEM.

12A-90.

DESCRIPTION.

12A-91.
12-90.

TROUBLE SHOOTING.

Refer to Section

Refer to paragraph

Refer to paragraph 12-89.
Refer to paragraph

12A-92. PRIMARY MAINTENANCE.
graph 12-91.
12A-93.

STARTER

Refer to para-

MOTOR.

12A-94. REMOVAL AND INSTALLATION.
a. Remove cowling in accordance with paragraph
12-3.
b. Remove induction airbox in accordance with
paragraph 12A-66.
c. Disconnect electrical power cable at starter
and insulate terminal as a safety precaution.
d. Remove nuts securing starter and remove
starter.
e. Reverse the preceding steps for reinstallation.
Install a new O-ring and be sure the starter drive
engages with the drive in the adapter.
12A-95.
12A-3.

EXHAUST SYSTEM.

Refer to figure

12A-96. DESCRIPTION. The exhaust system consists of two exhaust stack assemblies, one for the
left and one for the right bank of cylinders. These
exhaust stack assemblies are joined together to
route the exhaust from all cylinders through the
waste-gate or turbine. The three risers on the
left bank of cylinders are joined together into a
common pipe to form the left stack assembly. The
right rear cylinder exhaust is routed down and aft
to the rear of the engine where it connects to the
left stack assembly. The risers on the two right
front cylinders are connected to a common pipe to
form the right stack assembly. The right stack
assembly connects to the left stack assembly at
the front of the engine. Mounting pads for the
waste-gate and turbine are provided on the right
stack assembly. From the exhaust port of the turbine, a tailpipe routes the exhaust overboard through
the lower fuselage. The exhaust port of the wastegate is routed into the tailpipe so the exhaust gas can
be expelled from the system when not needed at the
turbine. The waste-gate is actuated by the wastegate actuator which, in turn, is controlled by the
waste-gate controller. Also, sleeving is installed
on the fuel hose from the engine-driven pump to the
fuel metering body and on the hose from the auxiliary
fuel pump to the engine-driven pump. This is to prevent excessive heat on these fuel hoses as they route
close to the exhaust stack.
12A-97. REMOVAL.
a. Remove engine cowling and right and left nose
caps in accordance with paragraph 12-3.

b. Remove intake manifold balance tube from front
of engine.
c. Remove heat shield at front of engine.
d. Loosen clamp and disconnect flexible duct at aft
end of cabin heater shroud on left exhaust stack
assembly.
e. Remove clamps and bolts securing rear heat
shield to engine and remove heat shield.
f. Remove clamps attaching left exhaust stack
assembly to riser pipes and to rear crossover pipe
on left side of engine.
g. Work left exhaust stack assembly down from
risers and out of crossover pipes at front and rear
of engine.
h. Remove four nuts and washers attaching exhaust riser pipe to each cylinder on left bank of cylinders and remove riser pipes and gaskets.
i. Remove clamp attaching exhaust tailpipe to exhaust port of turbine.
j. Remove bolts attaching waste-gate to right exhaust stack assembly. Work tailpipe from turbine
and lower waste-gate and tailpipe into cowling.
k. Remove bolts attaching turbocharger to mounting brackets.
1. Remove bolts and nuts attaching turbocharger
to right exhaust stack assembly. Lower turbocharger
into cowling.
m. Remove bolts, nuts and clamps attaching right
exhaust stack assembly to riser pipes on right side
of engine.
n. Work right exhaust stack assembly down from
risers and remove.
o. Remove nuts and washers attaching riser pipes
to front two cylinders on right side of engine and
remove riser pipes and gaskets.
p. Remove nuts and washers attaching exhaust pipe
to rear cylinder on right side of engine and remove
pipe and gasket.
12A-98.

INSTALLATION.
NOTE

It is important that the complete exhaust system, including the turbocharger and wastegate, be installed without pre-loading any
section of the exhaust stack assembly.
a. Use new gaskets between exhaust stacks and engine cylinders, at each end of waste-gate and between
turbocharger and exhaust stack.
b. Place all sections of exhaust stacks in position
and torque nuts attaching them to the cylinders evenly
to 100-110 lb-in., while riser clamps are loose.
c. Manually check that crossover pipe slip-joints do
not bind. Tighten clamp attaching left risers to left
stack assembly. Tighten the clamp attaching right
stack to right front riser.
d. Raise turbocharger into position and install bolts
and nuts attaching turbocharger to right exhaust stack
and those attaching turbocharger to front and rear
turbocharger supports (figure 12A-5). Tighten bolts
securely.

Change 1

12A-17

INTAKE
ATTACHES
TO CYLINDERS

ATTACHES
TO ENGINE
HEAT
SHIELD

INTAKE

TURBINE
INSTALLED

-

i
A

HERE
i|?
TAILPIPE

Figure 12A-3.
12A-18

Change I

V-WASTE GATE
INSTALLED
HERE

Exhaust System (Sheet 1 of 2)

4

1.
2.
3.
4.
5.

Clamp
Crankcase
Intake Manifold Balance Tube
Heat Deflector
Rivet

Figure 12A-3.

6.
7.
8.
9.

Heat Shield
Bolt
Lockwasher
Washer

10. Right Nosecap
11. Insulation
12. Retaining Skin
13. Left Nosecap
14. Screw

Exhaust System (Sheet 2 of 2)
12A-19
12A-19

e. Install bolts and nuts attaching waste-gate to
right hand exhaust stack and tighten securely.
f. While applying an upward force of one G to
counteract weight of turbocharger and waste-gate
assembly, tighten clamp attaching exhaust stack to
riser.
g. Tighten clamp securing tailpipe to turbocharger.
h. Be sure all parts are secure and safetled as re-

quired, then perform step "b" of paragraph 12A-99
to check for air leaks.
i. Install heater shroud duct and heat shields.
j. Install intake manifold balance tube at front of
engine and install heat shields at front of engine,
then install nose caps and cowling.
NOTE
The lower sections of turbocharger supports (index 8, figure 12A-5) are supplied as service parts
with their upper holes omitted. These undrilled
parts are also supplied when a new turbocharger
inlet stack, right front stack, or either of the
two right front risers is ordered. The following steps outline the proper procedure for
drilling and installing the supports.

should be made to detect cracks causing leaks which
could result in loss of optimum turbocharger efficiency and engine power. To inspect the engine exhaust system, proceed as follows:
a. Remove engine cowling as required so that ALL
surfaces of the exhaust assemblies can be visually
inspected.

WARNING
Never use highly flammable solvents
on engine exhaust systems. Never use
a wire brush or abrasives to clean exhaust systems or mark on the system
with lead pencils.

NOTE
Especially check the areas adjacent to welds
and slip joints. Look for gas deposits in surrounding areas, indicating that exhaust gases
are escaping through a crack or hole or

around the slip joints.
k. Install all parts but do not tighten attaching
clamps or bolts.
1. Torque nuts attaching risers to cylinders evenly
to 100-110 lb-in.
m. Tighten bolts and clamps per steps "d" through
"g".
NOTE
It is important that weight of turbocharger and
waste-gate assembly be counteracted, as listed
in step "f", when tightening clamps attaching
stacks to risers.
n. Make hole locations in undrilled supports to
match existing holes in upper supports.
o. Remove lower supports, leaving all other parts
tight.
-p. Drill the marked holes with a 3/8-inch drill.
On earlier models the holes were 0. 257-Inch, therefore, it may be necessary to enlarge the holes in
upper supports.
q. Reinstall supports, install bolts fastening upper
and lower supports together, then tighten all bolts
securely. If any exhaust system bolts or clamps
were loosened while lower supports were not installed, loosen all clamps and bolts and repeat the installation procedure to be sure no pre-loading is present.
r. Be sure all parts are secure and safetled as required, reinstall any parts removed for access, then
install nose caps and cowling.
12A-99. INSPECTION. Since exhaust systems of
this type are subject to burning, crackingand general
deterioration from alternate thermal stresses and
vibrations, inspection is important and should be accomplished every 50 hours of operation. Also, a
thorough inspection of the engine exhaust system

12A-20

Change 3

b. After visual inspection, an air leak check should
be made on the exhaust system as follows:
1. Attach the pressure side of an industrial
vacuum cleaner to the tailpipe opening, using a rubber plug to effect a seal as required.
NOTE
The inside of the vacuum cleaner hose should
be free of any contamination that might be
blown into the engine exhaust system.
2. With vacuum cleaner operating, all joints
in the exhaust system may be checked manually by
feel, or by using a soap and water solution and
watching for bubbles. Forming of bubbles is acceptable, if bubbles are blown away system is not
acceptable. Also, some bubbles will appear at the
joint of the turbocharger turbine and compressor
bearing housing.
c. Where a surface is not accessible for a visual
inspection, or for a more positive test, the following
procedure is recommended.
1. Remove exhaust stack assemblies.
2. Use rubber expansion plugs to seal openings.
3. Using a manometer or gage, apply approximately 1-1/2 psi (3 inches of mercury) air pressure
while each stack assembly is submerged in water.
Any leaks will appear as bubbles and can be readily
detected.
4. It is recommended that exhaust stacks found
defective be replaced before the next flight.
d. After installation of exhaust system components
perform the inspection in step "b" of this paragraph
to ascertain that sysyem is acceptable.

D2007C3-13 Temporary Change 2
22 February 1978

12A-100.

TURBOCHARGER.

12A-101. DESCRIPTION. The turbocharger is an
exhaust gas-driven compressor, or air pump, which
provides high velocity air to the engine intake manil.
fold. The turbocharger is composed of a turbine
wheel, compressor wheel, turbine housing and comw
opressor
housing.
presser
housing. The
The turbine,
turbine, compressor
compressor wheel
wheel
and interconnecting drive shaft comprise one com-et
plete assembly and are the only moving parts in-13
the turbocharger. Turbocharger bearings are lubricated with filtered oil supplied from the engine oil
system. Engine exhaust gas enters the turbine
housing to drive the turbine wheel. The turbine
wheel, in turn, drives the compressor wheel, producing a high velocity of air entering the engine indumped
gas is
is then
Exhaust gas
manifold. Exhaust
intake manifold.
duction intake
duction
then dumped
overboard through the exhaust outlet of the turbine
housing and exhaust tailpipe. Air is drawn into the
through
air
compressor
compressor through the
the induction
induction
air filter
filter and
and is
is
forced out of the compressor housing through a
tangential outlet to the intake manifold. The degree
of turbocharging is varied by means of a waste-gate
valve, which varies the amount of exhaust gas allowed
to~the
.. bypass
the turbine
to bypass the turbine.
12A-102. REMOVAL AND INSTALLATION.
a. Remove engine cowling as required.
b. Remove waste-gate to tailpipe clamp,
c. Loosen clamp at turbine exhaust outlet and work
tailpipe from turbine outlet.
d. Loosen clamps and remove air inlet and outlet
ducts from turbocharger compressor.
e. Disconnect oil pressure and scavenger lines
from turbocharger. Plug or cap open oil lines and
fittings. Remove clamp on oil supply line to the
turbocharger.
1. Loosen clamp and remove induction air inlet
elbow at turbocharger compressor.
g. Remove right cowl flap by disconnecting control
at cowl flap and removing hinge pin.
h. Cut safety wire and remove two bolts attaching
turbine to forward mounting bracket.
i. Remove three bolts attaching turbine to turbine
rear mounting bracket.

j. Remove three remaining bolts, washers and
nuts attaching turbine to exhaust manifold.
k.
k. Work
Work turbocharger
turbocharger from
from aircraft
aircraft through
through cowl
cowl
flap opening in lower cowling
Reverse the preceding steps for reinstallation.
When installing the turbocharger, install a new gasket between exhaust manifold and turbine exhaust

AC

ta saety wire.
ONTRO
R AN
AOWASTE-GATE

12A-1.

NCTIONS.

WA

AT

The waste-gate actuator

and controller uses engine oil for power supply. The
turbocharger
wast
controlled by
by the
the waste-gate,
wastegate, wasteturbocharger iss controlled
gate actuator, the absolute pressure and overboost
control
valve.
The
ctr
va
he waste-gate
te-gate bypasses engine
ine exe
haust gas around the turbocharger turbine inlet.
The waste-gate actuator, which is physically connected to the waste-gate by mechanical linkage, controls
trols the
the position
positon of
of the
the waste-gate
wate-gate butterfly
butterfly valve.
valve
The
solute pressure controller controls the maxicompressor discharge pressure,
mum turbocharger
rboos coro
e pre
s an
essive
overboost control valve prevents an excessive
pressure increase from the turbocharger compressor.
12A-105. OPERATION. The waste-gate actuator is
spring-loaded to position the waste-gate to the normally open position when there is not adequate oil
pressure in the waste-gate actuator power cylinder
during engine shut down. When the engine is started,
oil pressure is fed into the waste-gate actuator power
cylinder through the capillary tube. This automatically fills the waste-gate actuator power cylinder and
lines leading to the controllers, blocking the flow of
oil by normally closed metering and/or poppet valves.
As oil pressure builds up in the waste-gate actuator
power cylinder, it overcomes the force of the wastegate open spring, closing the waste-gate. When the
waste-gate begins to close, the exhaust gases are
routed through the turbocharger turbine. As the engine increases its power and speed, the increase of

SHOP NOTES:

Change 1

12A-21

TO TURBINE (WASTE GATE CLOSED)
OVERBOARD (WASTE GATE OPEN)
WASTE-GATE
ACTUATOR
(SPRING-LOADED

NOZZLES
... PUMP
- \8 |

A1R FIGATE I

/

RINDUCTI

INDUCT IOIn

IDCTONE
S
INDUCTION

TOFUl

DISCHARGE
.TO
FUELILTE
FUEL
TO
TO
FUEL
THROTTLE
TOST
I |T| OVDISCHARGE
| I
r'-|

AIRILTER
EXHAUST
ALTERNATE
-]1r

.
,
YASTE-ATE
FLO
|MNIFOLD
W THR
U

VALVEI.

I

|J

12A -22 CONTROL

////

.

(CLOSED BY
MAGNET)

R|1
AIR OVERBOARD
N
LR PRAM
CoE OMPRESSED
AIR

\\\ - MECHANICMPSSL LINKAGE

/-

FT
TO

AGAGE

I

^J

OIN

S

UMP
OIL PIL

LEGEND:

*~~~~~~~WASTE-GATE
^

EXHAUST AIRL

'.'
----

Figure 12A-4.
12A-22

Turbocharger System Schematic

ENGINE OIL
MECHANICAL LINKAGE

temperature and pressure of the exhaust gases causes
the turbocharger to rotate faster, raising the turbocharger compressor outlet pressure. As the compressor outlet pressure rises, the aneroid bellows
and the absolute pressure controller sense the increase in pressure. When at high engine speed and
load and the proper absolute pressure is reached, the
force on the aneroid bellows opens the normally
closed metering valve. When the oil pressure in the
waste-gate actuator power cylinder is lowered sufficiently, the waste-gate actuator open spring forces
the mechanical linkage to open the waste-gate. A
portion of the exhaust gases then bypasses the turbocharger turbine, thus preventing further increase of
turbocharger speed and holding the compressor discharge absolute pressure to the desired valve. Con-

versely, at engine idle, the turbocharger runs slowly
with low compressor pressure output; therefore, the
low pressure applied to aneroid bellows is not sufficient to affect the unseating of the normally closed
metering valve. Consequently, engine oil pressure
keeps the waste-gate closed. The overboost control
valve acts as a pressure relief valve and will open to
prevent an excessive pressure increase from the
turbocharger compressor. Above 19,000 feet, the
absolute pressure controller will continue to maintain
32. 5±. 5 inches of mercury manifold pressure at full
throttle. It is necessary to reduce manifold pressure
with the throttle to follow the maximum manifold
pressure versus altitude schedule shown on the instrument panel placard.

CAUTION
All turbocharged engine installations on Cessna aircraft are equipped with controller systems
which automatically control the engine within prescribed manifold pressure limits. Although
these automatic controller systems are very reliable and eliminate the need for manual control
through constant throttle manipulation, they are not infallible. For instance, such things as
rapid throttle manipulation (especially with cold oil), momentary waste-gate sticking, air in
the oil system of the controller, etc., can cause overboosting.
Consequently, it is still necessary that the pilot observe and be prepared to control the manifold pressure, particularly during take-off and power changes in flight.
The slight overboosting of manifold pressure beyond established minimums, which is occasionally
experienced during initial take-off roll or during a change to full throttle operation in flight, is
not considered detrimental to the engine as long as it is momentary. Momentary overboost is
generally in the area of 2 to 3 inches and can usually be controlled by slower throttle movement.
No corrective action is required where momentary overboosting corrects itself and is followed
by normal engine operation. However, if overboosting of this nature persists, or if the amount
of overboost goes as high as 6 inches, the controller and overboost control should be checked
for necessary adjustment or replacement of the malfunctioning component.
OVERBOOST EXCEEDING 6 INCHES beyond established minimums is excessive and can result
in engine damage. It is recommended that overboosting of this nature be reported to your
Cessna Dealer, who will be glad to determine what, if any, corrective action needs to be taken.

12A -23

12A- 106.

TROUBLE SHOOTING.

TROUBLE
UNABLE TO GET RATED
POWER BECAUSE MANIFOLD PRESSURE IS LOW.

REMEDY

PROBABLE CAUSE
Controller not getting enough oil
pressure to close thewaste-gate.

Check oil pump outlet pressure, oil
filter and external lines for obstructions. Clean lines and replace if defective. Replace oil
filter.

Controller out of adjustment or
defective.

Refer to paragraph 12A-109.
Replace controller if defective.

Defective actuator.

Refer to paragraph 12A-111.
place actuator if defective.

Leak in exhaust system.

Check for cracks and other obvious defects. Replace defective
components. Tighten clamps and
connections.

Leak in intake system.

Check for cracks and loose
connections. Replace defective
components. Tighten all clamps
and connections.

Defective controller.

Refer to paragraph 12A-109.
Replace if not adjustable.

Waste-gate actuator linkage
binding.

Refer to paragraph 12A-111.

Waste-gate actuator leaking
oil.

Replace actuator.

Turbocharger overspeeding from
defective or improperly adjusted
controller.

Refer to paragraph 12A-109.
Replace if defective.

Waste-gate sticking closed.

Correct cause of sticking. Refer
to paragraph 12A-109. Replace
defective parts.

Controller drain line (oil return
to engine sump) obstructed.

Clean line.

ENGINE POWER INCREASES
SLOWLY OR SEVERE MANIFOLD PRESSURE FLUCTUATIONS WHEN THROTTLE
ADVANCED RAPIDLY.

Overboost control valve out of
adjustment or defective.

Replace if defective.

Waste-gate operation is
sluggish.

Refer to paragraph 12A-111.
Replace if defective. Correct
cause of sluggish operation.

ENGINE POWER INCREASES
RAPIDLY AND MANIFOLD
PRESSURE OVERBOOSTS
WHEN THROTTLE ADVANCED RAPIDLY.

Overboost control valve out of
adjustment or defective.

Replace if defective.

Waste-gate operation is
sluggish.

Refer to paragraph 12A-111.
Replace if defective. Correct
cause of sluggish operation.

ENGINE SURGES OR
SMOKES.

TURBOCHARGER NOISY
WITH PLENTY OF POWER.

12A-24

Re-

_

Replace if defective.

12A-106.

TROUBLE SHOOTING (Cont).
TROUBLE

FUEL PRESSURE DECREASES
DURING CLIMB, WHILE MANIFOLD PRESSURE REMAINS
CONSTANT.

PROBABLE CAUSE

REMEDY

Compressor discharge pressure
line to fuel pump aneroid
restricted.

Check and clean out restrictions.

Leaking or otherwise defective
engine-driven fuel pump
aneroid.

Replace engine-driven
fuel pump.

Leak in intake system.

Check for cracks and other
obvious defects. Tighten all
hose clamps and fittings.
Replace defective components.

Leak in exhaust system.

Check for cracks and other
obvious defects. Tighten all
clamps and fittings. Replace
defective components.

Leak in compressor discharge
pressure line to controller.

Check for cracks and other
obvious defects. Tighten all
clamps and fittings. Replace
defective components.

Controller seal leaking.

Replace controller.

Waste-gate actuator leaking oil.

Replace actuator.

Waste-gate butterfly - closed gap
is excessive.

Refer to paragraph 12A-111.

Intake air filter obstructed.

Service air filter. Refer to
Section 2 for servicing
instructions.

Defective engine-driven fuel
pump aneroid mechanism.

Replace engine-driven fuel
pump.

Obstruction or leak in compressor
discharge pressure line to enginedriven fuel pump.

Check for leaks or obstruction.
Clean out lines and tighten
all connections.

FUEL FLOW INDICATOR
DOES NOT REGISTER
CHANGE IN POWER SETTINGS
AT HIGH ALTITUDES.

Moisture freezing in indicator
line.

Disconnect lines, thaw ice and
clean out lines.

SUDDEN POWER DECREASE
ACCOMPANIED BY LOUD
NOISE OF RUSHING AIR.

Intake system air leak from
hose becoming detached.

Check hose condition. Install
hose and hose clamp securely.

MANIFOLD PRESSURE GAGE
INDICATION WILL NOT REMAIN STEADY AT CONSTANT
POWER SETTINGS.

Defective controller.

Replace controller.

Waste-gate operation is
sluggish.

Refer to paragraph 12A-111.
Replace if defective. Correct
cause of sluggish operation.

MANIFOLD PRESSURE DECREASES DURING CLIMB
AT ALTITUDES BELOW NORMAL PART THROTTLE
CRITICAL ALTITUDE, OR
POOR TURBOCHARGER
PERFORMANCE
INDICATED BY CRUISE
RPM FOR CLOSED WASTEGATE. (Refer to paragraph
12A-107.)

FUEL FLOW DOES NOT DECREASE AS MANIFOLD
PRESSURE DECREASES AT
PART-THROTTLE
CRITICAL ALTITUDE.

12A-25

12A-107. CONTROLLER AND TURBOCHARGER OPERATIONAL FLIGHT CHECK. The following procedure details the method of checking the operation of the absolute controller overboost control valve, and a performance
check of the turbocharger.
1

TAKE-OFF-ABSOLUTE CONTROLLER CHECK.
a. Cowl Flaps - Open.
b. Airspeed - 110 MPH IAS.
c. Oil Temperature - Middle of green arc.
d. Engine Speed - 2700 * 25 RPM.
e. Fuel Flow - 28.0 to 29.5 GPH (168.0 to 177.0 LBS/HR) (Full Rich Mixture).
1. Full Throttle M. P. - Absolute controller should maintain 32.5 ± .5 in. Hg (stabilized).

Climb 2000 feet after take-off to be sure manifold pressure has stabilized. It is normal on the first take-off of
the day for full throttle manifold pressure to decrease 1/2 to 1.0 inch of mercury within one minute after the
initial application of full power. Refer to paragraph 12A-109 for absolute controller adjustment.
CLIMB - ABSOLUTE CONTROLLER AND TURBOCHARGER PERFORMANCE CHECK.
a. Cowl Flaps - Open.
b. Airspeed - 120 MPH IAS.
c. Engine Speed - 2500 RPM.
d. Fuel Flow - Adjust mixture for 20 GPH (120.0 LBS/HR).
e. Part - Throttle M. P. - 27.5 in. Hg.
f. Climb to 20,000 feet - Check part-throttle critical altitude during climb.

2

This part-throttle critical altitude is where manifold pressure starts decreasing during the climb at a rate of
approximately 1.0 inch of mercury per 1000 feet. After noting this altitude and the outside air temperature,
the desired manifold pressure should be maintained by advancing the throttle during the remainder of the climb.
Once the climb power setting is established after take-off, the controller should maintain a steady manifold
pressure up to the part-throttle critical altitude indicated in the following chart. If part-throttle critical
altitude has not been reached by 20, 000 feet, discontinue check and proceed to cruise check.
Outside Air Temperature

Part-Throttle Critical Altitude (75% Power)

Standard or Colder
20°F Above Standard
40'F Above Standard

Above 24,000 feet
16,000 to 22,000 feet
10,000 to 16,000 feet

Part-throttle critical altitudes lower than those listed indicate the turbocharger system is not operating
properly (refer to the trouble shooting chart in paragraph 12A-106). Critical altitudes above those listed
indicate turbocharger performance better than normal. Also check that fuel flow decreases as manifold
pressure decreases at critical altitude. Refer to the trouble shooting chart if fuel flow does not decrease.
3 CRUISE - TURBOCHARGER PERFORMANCE CHECK.
a. Cowl Flaps - Closed.
b. Airspeed - Level flight.
c. Pressure Altitude - 20,000 feet.
d. Engine Speed - 2700 RPM.
e. Part-Throttle M. P. - 27.5 in. Hg.
f. Fuel Flow - Lean to 18 GPH (108.0 LBS/HR).
g. Propeller Control (1) Slowly decrease RPM until manifold pressure starts to drop, indicating waste-gate is closed.
(2) Note outside air temperature and RPM as manifold pressure starts to drop, which should be in
accordance with the following chart.
(3) After noting temperature and RPM, increase engine speed 50 RPM to stabilize manifold pressure, with the waste-gate modulating exhaust flow to control compressor output.
Outside Air Temperature

RPM where M. P. Starts to Decrease

40°F Above Standard
20°F Above Standard
Standard Temperature
20°F Below Standard
40°F Below Standard

2700
2600
2500
2400
2300

to
to
to
to
to

2550
2450
2350
2550
2150

If the waste-gate is closed at engine speeds higher than those listed, refer to the trouble shooting chart in
paragraph 12A-106. Closing of the waste-gate at engine speeds lower than those listed indicates turbocharger
performance better than normal.
12A-26

20,000 FT
PRESSURE
ALTITUDE

2000 FT
ABOVE
GROUND
NOTE
Circled numbers refer to corresponding
flight checks required in preceding text.

12A-108. REMOVAL AND INSTALLATION OF TURBOCHARGER CONTROLLER.
a. Disconnect and tag oil lines from controller and
plug or cap open lines and fittings.
b. Disconnect compressor outlet pressure sensing
line from controller and plug or cap open line and
fitting.
c. Remove two bolts attaching controller to mounting bracket on firewall.
d. Remove controller from aircraft, being careful
not to drop controller unit.
e. Installation of the controller may be accomplished by reversing the preceding steps. Resafety bolts
attaching controller to bracket.
12A-109. ABSOLUTE CONTROLLER ADJUSTMENTS.
(Refer to figure 12A-6.)
a. With engine oil temperature at middle of green
arc, slowly open throttle and note maximum manifold
pressure obtainable. Do not exceed 32. 5±. 5 in. Hg.
b. Cut safety wire and remove plug from bottom of
absolute controller (the vertical unit).
c. Using a flat-bladed screwdriver, rotate metering valve seat clockwise to increase manifold pressure and counterclockwise to decrease manifold pressure. Lightly tap the unit after each adjustment to
seat internal parts.

NOTE
When adjusting, rotate in VERY small increments as this is an extremely sensitive
adjustment. Approximately 13 degrees rotation will change the manifold pressure
reading about one inch Hg.
d. Install and safety plug in absolute unit, then
operate engine as in step "a" to ascertain that adjustment has not caused radical change in manifold
pressure.
NOTE
When making adjustment on the ground, the
hotter the engine gets. the lower the manifold
pressure will be.
e. After each adjustment, the aircraft must be
flight tested to check results.
f. Repeat this procedure until desired results are
obtained.
12A-110. REMOVAL AND INSTALLATION OF
WASTE-GATE AND ACTUATOR.
a. Disconnect and tag oil lines from actuator and
plug or cap open lines and fittings.
12A-27

5

* Beginning with aircraft U20601605 and
on and all service parts, a new oil
inlet adapter (22) and check valve (11)
is used. If a new adapter or check
valve is installed on aircraft prior to
U20601604, it will be necessary to install both the check valve and oil inlet

6

4

adapter.

-

TO
WITH
FUEL INJECTION,
VACUUM
WITHOUT
SYSTEM
PRESSURE
IAIR
1
SYSTEM
VACUUM
SYSTEM
-

INTAKE
SYSTEMtC

TO
SCAVENGER/
PUMP
ENGINE

TO

TO

FROM
AIR INLET

SYSTEM

Safety wire these

/14

/

position.
is within60

°

of top

Whenever bolts attaching oil
outlet adapter (18) or inlet

adapter (22) are loosened or
sure
that
be lockremoved,

washers are in good condiion, torque the bolts
bolts torque
to 180
to 190 lb-in., then safety the
outboard bolts together and

16

Figure 12A-5.
12A-28

Change 1

the inboard bolts together.

Turbocharger System (Sheet 1 of 2)

24 25

26

7

8

Beginning with aircraft serial
the duct supports

0
2U20602568

'0

'-/2

* NOTE

have slotted holes or adjust-

*

31

1. Turbocharger Discharge Duct
2. Coupler
3. O-Ring Seal
4. Overboost Control Valve
5. Throttle Body Adapter
6. Absolute Controller
7. Vacuum System Oil Separator
8. Turbocharger Support
9. Washer
10. Cotter Pin
11. Check Valve
12. Tube
13. Shaft
14. Exhaust System
15. Waste-Gate Actuator
16. Waste-Gate
17. Tail Pipe
18. Oil Outlet Adapter
Figure 12A-5.

19.
20.
21.
22.
23.
24.
25,
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.

Check Valve
Turbine Housing
Compressor Housing
Oil Inlet Adapter
Turbocharger Inlet Hose
Doubler
Magnet
Door
Induction Air Filter
Baffle Assembly
Center Duct
Forward Duct
Seal
Wire Guard
Nose Cap Flexible Duct
Turbocharger Inlet Elbow
Aft Duct
Support

Turbocharger System (Sheet 2 of 2)
Change 3

12A-29

b. Remove bolts, washers and nuts attaching
waste-gate and actuator assembly to tailpipe.
c. Loosen clamp attaching tailpipe to turbine exhaust outlet and work tailpipe from turbine.
d. Remove bolts, washers and nuts attaching the
assembly to the exhaust manifold.
e. Remove the assembly from aircraft, being careful not to drop the unit.
f. Installation may be accomplished by reversing
the preceding steps.

ABSOLUTE PRESSURE
CONTROLLER

NOTE
When installing the assembly, be sure the
gaskets at inlet and outlet of valve are installed and are in good condition. Replace
gaskets if damaged.
12A-111. ADJUSTMENT OF WASTE-GATE ACTUATOR. (Refer to figure 12A-7.)
a. Remove waste-gate actuator in accordance with
paragraph 12A-110.
b. Plug actuator outlet port and apply a 50 to 60
psig air pressure to the inlet port of the actuator.
c. Check for 0.010 + 0-.005 inch gap between butterfly and waste-gate body as shown in figure 12A-7.
d. If adjustment is required, remove pin from
actuator shaft.
e. Hold clevis end and turn shaft clockwise to increase gap or counterclockwise to decrease gap of
butterfly. Install pin through clevis and shaft, securing pin with washer and cotter pin.
f. After adjusting closed position and with zero
pressure in cylinder, check butterfly for a clearance
of 1. 100 + .000 -. 125 inch in the full-open position
as shown in figure 12A-7.
g. If adjustment is required, loosen locknut and
turn stop screw clockwise to decrease or counterclockwise to increase clearance of butterfly.
h. Recheck butterfly in the closed position to ascertain that gap tolerance has been maintained.
NOTE
To assure correct spring loads, actuate
butterfly with air pressure. Actuator shaft
and butterfly should move freely. Actuator
shaft should start to move at 15±2 psig and
fully extend at 35±2 psig. Two to four psi
hysteresis is normal, due to friction of Oring against cylinder wall.
i. Remove air pressure line and plug from actuator.
j. Install waste-gate and actuator as outlined in
paragraph 12A-110.

12A -30

FLAT-BLADED SCEWDRIVER

Figure 12A-6.

Controller Adjustment

12A-112. EXTREME WEATHER MAINTENANCE.
Refer to paragraph 12-98.
12A-113.
12-99.

COLD WEATHER.

Refer to paragraph

12A-114. HOT WEATHER. When the engine is hot or
the outside air temperature is high, the engine may
die after running several seconds because the mixture
became either too lean due to fuel vapor or too rich
due to excessive prime fuel. The following procedure will prevent over-priming and take care of
fuel vapor in the system.
a. Set the throttle 1/3 to 1/2 open.
b. When the ignition key is on BOTH and you are
ready to engage the starter, turn the fuel pump on
HI until the fuel flow comes up to 4-6 gal/hr and
then turn the pump off.

1. 100+.000 -. 125

.010+.0-. 005

OUTLET

LOCKNUT

c.

12A-115. SEACOASTAND
HUMIDAREAS
paragraph
12-101. TO
TORQUE

Without hesitation, engage the starter and the
CLEVIS
CLEVIS END
END

PIN
PIN

Figure 12A-7.

Referto

SHAFT
SHAFT

Waste-Gate Adjustment

start should
in 3 to revolutions.Gate
Figurengine
Adjustment
12ANOTE
During a restart
adown
briefand
shut-down
in
engine after
will
slowgradually
stop. When
gineextremely
speed start
ater
a the
brief
shut-downn
weather,
hot weather,
presence
of fuel
vapor may require the pump to run on HI for
up to i1 minute or more before the vapor is
cleared sufficiently to obtain 4-6 gal/hr for
If is fuel
vapor
NOd.
it will
into starting.
thethere
injector
nozzles
in in2 to 3 seconds
and pass
the

por. Without hesitatient
use of Hengine
boostarter
aneeded
sithethe
pro
revolutions.
should
startin 3 to Adjust
for 1200-1400 RPM.
throttle for
one second
to clear
out
the
vad. If there is for
fuel approximately
vapor
in the lines,
it will pass
into the injector nozzles in 2 to 3 seconds and the
engine will gradually slow down and stop. When engine speed starts to decrease, turn the fuel pump on
HI for approximately one second to clear out the vapor. Intermittent use of HI boost is needed since pro-

longed use of HI pump after the vapor is cleared will
flood out the engine.
e.
Let the engine
at 1200 to 1400 RPM until the
ento
paragraph
12-103ne. run
e.
Lthe
engine
run
to 1400
until the
vapor
is eliminated
and at
the 200
engine
idlesRPM
normally.
If prolonged cranking is necessary, allow the starter
motor to cool at frequent intervals, since excessive
heat may damage the armature.
s
12A-117.leareds,
GROUND SERVICE RECEPTACLE.
Refer

paragraph
12-101.
A
AREAS. Refer to paragraph 12-102.
12A-116. DUSTY AREAS.
12A-115. SEA
AANKND HUMID AR EAS. Refer toa
12A-117. GROUND SERVICE RECEPTACLE. Refer
to paragraph 12-103.
12A-118.
12-104.

HAND CRANKING.

Refer to paragraph

12A-31

AIR MANIFOLD

EXISTING HOSE TO FUEL PUMP GAGE VENT

EXISTING
ELBOW

FUEL METERING
UNIT

FUEL PRESSURE GAGE
EXISTING
ELBOW

4

EXISTING FUEL PUMP
OUTLET HOSE

3

ENGINE DRIVEN
FUEL PUMP

1.
2.
3.
4.
5.

Test Hose Assembly (S-1168-3-96)
Test Hose Assembly (S-1168-4-8. 5)
Nipple (AN816-3D)
Tee (AN917-1D)
Nipple (AN816-3D)

NOTE
When adjusting the fuel injection pump unmetered fuel pressure, the test equipment
may be "teed" into the engine driven fuel pump outlet hose at the fuel metering unit
and to the existing elbow on the air manifold.
Figure 12A-8.
12A-32

Fuel Injection Pump Adjustment Test Harness (Turbocharged Engine)
D2007C3-13 Temporary Change 2
22 February 1978

SECTION 13
FUEL SYSTEM
TABLE OF CONTENTS

Page

FUEL SYSTEM
..............
Description ..............
Precautions ..........
Trouble Shooting . . . . ...
.
Fuel Vents.3
........
Description .
..........
Checking .............
Fuel Cells .............
Description .
..........
General Precautions .........
Removal
.............
Repair ...........
Installation .......
Fuel Quantity Transmitters
Description ........
Removal and Installation ......
13-1.

.
..

.
.

13-1
13-1
. 13..
. 13-2
. 13-9
. 13-9
13-9
. 13-9
. 13-9
13-9
13-9
13-12
...
13-12
......13-12
..
13-12
13-12

FUEL SYSTEM.
NOTE

The fuel system as described in this section
does not include the fuel injection system.
Refer to Section 12 or 12A for that part of
the fuel system.
13-2. DESCRIPTION. Fuel from the cells in the
wings is gravity-fed through fuel reservoir tanks installed forward of the front doorpost bulkheads, beneath the cabin floor, to the engine driven fuel pump
The fuel line from the lower forward corner of each
fuel cell to the reservoir tank serves as a combination fuel feed and vapor return line. The fuel bypasses the electric auxiliary fuel pump when the pump is
not in operation. The fuel cells are individually vented overboard through check valves located in each
cell.
13-3.

Removal and Installation of Fuel
Reservoir Tanks .....
. . ..... 13-12
Removal and Installation of Fuel
Selector Valve ....
..
..
. . . .13 15
Fuel Seletor Valve Repair .
.
13-15
Auxiliary Electric Fuel Pump ..
13-15
Removal and Installation ...
13-16
Electric Fuel Pump Circuits
.
13-17
Rigging Throttle-Operated Switch
13 18
Fuel Flow Test ........
.13-19
Maximum High Boost Check ..
. .13-191
Fuel Strainer .............
13-19
Disassembly and Assembly .....
13-19
Electric Fuel Quantity Indicators
and Transmitters
.........
13-19
a. During all fueling. defueling, tank purging, and
tank repairing or disassembly, ground the airplane to
a suitable ground stake.
b. Residual fuel draining from lines and hoses constitutes a fire hazard. Use caution to prevent the
accumulation of fuel when lines or hoses are disconnected.
c. Cap open lines and cover connections to prevent
thread damage and the entrance of foreign matter

PRECAUTIONS.
NOTE

There are certain general precautions and
rules concerning the fuel system which
should be observed when performing the
operations and procedures in this Section.
These are as follows:

13-4.

NOTE
Throughout the aircraft fuel system, from
the fuel cells to the engine-driven fuel
pump, use NS-40 (RAS-4) (Snap-On Tools
Corp., Kenosha, Wisconsion), MIL-T-5544
(Thread Compound, Antiseize, Graphite
Petrolatum), USP Petrolatum, or engine oil
as a thread lubricant or to seal a leaking connection. Apply sparingly to male threads
only, omitting the first two threads, exercising extreme caution to avoid "stringing"
sealer across the end of the fitting. Always
ensure that a compound, the residue from a
previously used compound, or any other foreign material cannot enter the system.
Throughout the fuel injection system, from
the engine-driven fuel pump through the
discharge nozzles, use only a fuel-soluble
lubricant, such as engine oil, on fitting
threads. Do not use any other form of
thread compound on the injection system.

TROUBLE SHOOTING.
Use this chart in conjunction with the engine trouble shooting charts in Sections 12 and 12A.
TROUBLE

NO FUEL FLOW TO
ENGINE-DRIVEN
FUEL PUMP.

PROBABLE CAUSE

REMEDY

Fuel selector valve not turned on.

Turn fuel selector valve on.

Fuel cells empty.

Service with proper grade and
amount of fuel.

Change 3

13-1

13-4.

TROUBLE SHOOTING (Cont).
TROUBLE

NO FUEL FLOW TO
ENGINE-DRIVEN
FUEL PUMP. (Cont).

FUEL STARVATION
AFTER STARTING.

NO FUEL FLOW WHEN
ELECTRIC PUMP
OPERATED.

NO FUEL QUANTITY
INDICATION.

FLUCTUATING FUEL
PRESSURE INDICATIONS. (TURBO AIRCRAFT)

13-2

Change 2

REMEDY

PROBABLE CAUSE
Fuel line disconnected or broken.

Connect or repair fuel lines.

Fuel cell screen plugged.

Remove and clean screen.
Flush out fuel cell.

Defective fuel selector valve.

Remove and repair or replace
selector valve.

Plugged fuel strainer.

Remove and clean strainer and screen.

Defective check valve in electric
fuel pump.

Repair or replace electric pump.

Fuel line plugged.

Disconnect lines as necessary to
locate obstructions, then clean.

Partial fuel flow from the preceding causes.

Use the preceding remedies.

Malfunction of engine-driven fuel
pump or fuel injection system.

Refer to Section 12 or 12A.

Fuel vents plugged.

See paragraph 13-7.

Water in fuel.

Drain fuel tank sumps, fuel
lines, and fuel strainer.

Defective fuel pump switch.

Replace defective switch.

Open or defective circuit breaker.

Reset.

Loose connections or open
circuit.

Tighten connections; repair or
replace wiring.

Defective electric fuel pump.

Replace defective pump.

Defective engine-driven fuel
pump bypass or defective fuel
injection system.

Refer to Section 12 or 12A.

Fuel cells empty.

Service with proper grade and
amount of fuel.

Circuit breaker open or defective.

Reset.

Loose connections or open circuit.

Tighten connections; repair wiring.

Defective fuel quantity indicator.

Replace indicator or sending unit.

Obstructed filter in fuel inlet
strainer of metering unit.

Remove and clean.

Manifold valve.

Replace.

Fuel flow Indicator.

Replace.

Replace if defective.

Replace if defective.

FUEL
LEFT
FUEL TANK

INDICATORS

QUANTITY

LEFT

RIGHT

FUEL

TANK

FILLER
CAP

FILLER
CAP
FUEL

OUANTITY TRANSMITTERS

VENTENT
ASCREEN
DIN
VALVE

SCREEN

T

SELECTOR
VALVE

RAIN
VALVE
UON
DID

1

"

RESERVOIR
-"UEL
WITH DRAIN PLUG

'

FUEL RESERVOIR
WITH DRAIN PLUG

\

SOLENOID

CHECK

ALVE

VAjV
LVE (OPT

OIL

STRAINE

DILUTION
-AUX

DRAIN

TO OIL SYSTEM

THROTTr

BUS BAR -3

C
\

' <

//

~
P-8

*BUS.ARI<
\

\
\

IGNITION.
S7ARTER
SWITCH
\
\

j

FUEL
DISTRIBUTO

THRU 1970 MODELS

DE
FEL

ELECTICAL
CONNECTION

UNIT

L

GAGE
OW

|I

?~

)

||

NON.TURBOCHARGED

TURBOCHARGED

BOTH BAYS SIMULTANEOUSLY

NOZZLES

NOTE

FUEL
PRETURN

LINKAGE

ME.

|f:

1 FUEL SUPPLY
EXCESS FUEL
AND VAPOR

_^B

I---gF

D
FUEL

3
THROTTLE

:Il11111111[
I[IIIF11:1'1lI

MIXTURE

Vi

AC R EEEE

.... __MECHANICAL

S I1::',1111111111111:11

R

~\
.~..E*SCREEN
AN

PRIMER

ENGINE
FUEL
PUMP

FFILJTELR X»H

TO

ENGINE

|

OFF
FUEL PUMP
SWITCH

'\
THROTTlE
MCNA

£ENGINE

.

-

E \

,['~'

KNOB

This schematic shows the aircraft fuel system. For engine fuel injection
schematics, refer to Section 12 for the non-pressurized system used on
all non-turbocharged aircraft. Refer to Section 12A for the pressurized
system used on lurbocharged aircraft.

Figure 13-1.

Fuel System Schematic (Sheet 1 of 2)
13-3

FUEL QUANTITY INDICATORS

:

LEFT

RIGHT

LEFT FUEL CELL

RIGHT FUEL CELL

VENTED FILLER
FILLER CAP
CAP

VENTED

FUEL QUANTITY TRANSMITTERS

SELECTOR

VALVE

SCREEN

DRAIN
DR
VALVE

VALVE

SCREEN
VENT

VENT
FUEL RESERVOIR
WITH DRAIN PLUG

FUEL RESERVOIR
W
ITH DRAIN PLUG

I

1

FUEL PUMP

THROTTLE SWITCH

FUEL PUMP

FROM BUS BAR
THRU U20602199

r

--

HIc)CHECK

»-L

Lo

'

J
-'^

:

BEGINNING
WITH U20602200

:: : :

ENGI~NE FUEL PUMP
EN__NE FE
MIXTURE CONTROL

f

-

-L

f

AIR THROTTLE
I|

OT

(
FL

tE
0

:
EL :UNIT"/:

FUEL DISTRIBUTOR

HIt|O
G

l)

CODE

IR

|T

CHECK VALVE

, FROM STARTER SWITCH
STRAINER DRAIN KNOB
i FUEL STRAINER

SYS TS~EMpI
TOOIL VALVE (OPT)
SYSTEM
FILTER
THROTTLE

BU
...NGE
ENGINE PRIMER

'

OMMAX B
HOLD IN
OIL DILUTION
SWITCH (OPT)-

V

ENGINE A

i

I--

WSU^

I

g4

NNON-TURBOCHARGED
TURBOCHARGED
FLOW INDICATORS

4FUEL

TO CIRCUIT

BREAKER

FUEL NOZZLES

BEGINNING WITH 1971 MODELS
"'l::

FUEL SUPPLY
NOTE

EXCESS FUEL
AND VAPOR
RETURN FUEL
MECHANICAL
LINKAGE

This schematic shows the aircraft fuel system. For engine fuel injection schematlcs, refer to Section 1t for the non-pressurized system used on all nonturbocharged aircraft. Refer to Section 12A for the pressurized system used on
turbocharged aircraft.

Fuel pump switch cannot be shown in OFF position, schematically.
ELECTRICAL
CONNECTION

Fuel cannot be used from both fuel cells simultaneously.
Figure 13-1.

13-4

Change 2

Fuel System Schematic (Sheet 2 of 2)

REFER TO FIGURE 13-3

15

12.

1. Fuel Vent Line
2.
3.
4.
5.
6.
7.7.
8.
9.
9.
10
11.
10.
11.

Selector-to-Strainer Line

13. Vapor Return-to-Selector Line
Fuel Vent Valve
14. Electric Pump Drain Line
Forward Line Screen
15. Electric Fuel Pump
Aft Line Screen
Strainer
Aft Fuel Line
Figune Screen
13-2. Fuel System (Sheet 16.of Fuel
2)
Forward Fuel Line
Strainer
Drain Line
Line
Fuel Strainer
17. Fuel
Line
Return Check Valve
Aft
18. Vapor Return
Line e
Aft Fuel
Fuel Lin
Reservoir
19. Right Reservoir
LLeft Reservoir
eft
Selector
Line Valve
20.
Selector
Valve
Handle
Valve Line
Reservoir-to-Selector
Fuel Line
Forward
21.
Line
Slector-to-Reservoir
Return
Vapor
Fe
Valve
Selector
21. Forward Fuel Line
Vapor
Return Selector-to-Reservoir Line
Selector Valve

Figure 13-2.

13-5

Fuel System (Sheet 1 of 2)
13-5

3

-

REFER TO FIGURE 13-6
REFER TO FIGURE 13-3
REFER TO FIGURE 13-7

1.
2.
3.
4.

Fuel Vent Line
Fuel Vent Valve
Forward Line Screen
Aft Line Screen

5.
6.
7.
8.
9.

Left Reservoir
Fuel Selector Valve
Auxiliary Pump Drain Line
Fuel Strainer
Strainer Drain Control

Figure 13-2.
13-6

10.
11.
12.
13.

Fuel System (Sheet 2 of 2)

Vapor Return Check Valve
Auxiliary Fuel Pump
Right Reservoir
Fuel Selector Handle

NON-TURBOCHARGED

TURBOCHARGED

1. Clamp

5. O-Ring

2.

6.

Duct

3. Shroud Half
4. Auxiliary Fuel Pump

Reducer

7. Pump Bracket
8. Elbow

9. Drain Line
10.

Grommet

11.
12.

Fuel Strainer
Fuel Hose

Figure 13-3. Electric Fuel Pump and Strainer Installation (Sheet 1 of 2)
13-7

7

3

1. Reducer
2. O-Ring
3. Auxiliary Fuel Pump

Figure 13-3.
13-8
13-8

4. Clamp
5. Bracket

6.
7.
8.

Elbow
Fuel Strainer
Fuel Hose

Electric Fuel Pump and Strainer Installation (Sheet 2 of 2)

VIEW LOOKING FORWARD

VIEW LOOKING INBOARD

3
4

4

OUTBD

3.50"

1.12"

.19"
NOTE

1. Wing
2. Vent
3. Strut
4. Fairing

DIMENSIONS MUST BE
WITHIN ±.03" TOLERANCE.

Figure 13-4.
13-5.

Fuel Vent Location

FUEL VENTS.

f. Any fuel vent found plugged or restricted must
be corrected prior to returning airplane to service.

13-6.
DESCRIPTION. A fuel vent line is installed
in the outboard end of each fuel cell. The vent line
extends overboard down through the lower wing skin.
The inboard end of the vent line extends into the fuel
cell, then is offset downward from cell upper surface.
A vent valve is installed on the inboard end of the vent
line inside the fuel cell.
13-7.
CHECKING FUEL VENT. Field experience
has demonstrated that fuel vents can become plugged,
with possible fuel starvation of the engine, or collapse
of fuel cells. Also, the bleed hole in the vent valve
assembly could possibly become plugged, allowing
pressure from expanding fuel to pressurize the cells.
The following procedure may be used to check the
vent and bleed hole in the valve assembly.
a. Attach a rubber tube to the end of the vent line
beneath one wing.
b. Turn off fuel selector valve.
c. Blow into tube to slightly pressurize the tank.
If air can be blown into tank, the vent line is open.
d. After tank is slightly pressurized, insert end
of rubber tube into a container full of water and
watch for a continuous stream of bubbles, which indicates the bleed hole in valve assembly is open and
relieving pressure.
first.
e. Repeat steps "a" through "d" for fuel vent
beneath opposite wing.
NOTE
Remember that a plugged vent line or bleed
hole can cause either fuel starvation and
collapsing of fuel cell or the pressurizing
of the cell by fuel expansion.

NOTE
The fuel vent line protruding beneath the wing
near the wing strut must be correctly aligned
to avoid possible icing of the vent tube. Dimensions are shown in figure 13-4.
13-8.

FUEL CELLS.

(RUBBERIZED. )

13-9 DESCRIPTION. Rubberized. bladder-type
fuel cells are installed in the inboard bay of each wing
panel. These cells are secured by fasteners to prevent collapse of the flexible cells.
13-10. GENERAL PRECAUTIONS. When storing
inspecting or handling rubberized, bladder-type fuel
cells, the following precautions should be adhered to:
a. Fold cells as smoothly and lightly as possible
with a minimum number of folds. Place protective
wadding between folds.
b. Wrap cell in moisture-proof paper and place in
a suitable container. Do not crowd cell in container.
Use wadding to prevent movement.
c. Stack boxed cells to allow access to oldest cells
Do not allow stacks to crush bottom boxes.
Leave cells in boxes until used.
d. Storage area must be cool, +30°F to +85 ° , and
free of exposure to sunlight, dirt and damage.
e. Used cells must be cleaned with soap and warm
water prior to storage. Dry and package as outlined
in the preceding steps.
f. Do not carry cells by fittings. Maintain original
cell contours or folds when refolding for boxing.
13-11. FUEL CELL REMOVAL.
a. Drain fuel from applicable cell.

(Pages 13-10 and 13-11 Deleted)

Change 3

13-9

NOTE
Prior to removal of cell, drain fuel, purge
with fresh air, and swab out to remove all
traces of fuel.

tion
of a cell for conditions noted in the preceding
steps.
d. Install fuel drain adapter and snap fasteners.
e. Check to ensure cell is warm enough to be flexible and fold as necessary to fit through fuel cell

access opening.
b. Remove wing root fairings and disconnect fuel
lines at wing root.
c. Remove clamps from forward and aft fuel cell
bosses at wing root and carefully work fuel strainers
and lines from cell bosses.
d. Disconnect electrical lead and ground strap from
fuel quantity transmitter and carefully work transmitter from fuel cell and wing rib.
e. Remove screws attaching drain adapter to lower
surface of wing.
f. Remove clamps attaching crossover vent line to
fuel cells and work vent line out of cell being removed.
In aircraft equipped with long-range cells, remove
vent extension tube from inside cell. Vent extension
tube is attached to the crossover vent bars on the cell.
g. Remove fuel filler adapter and gaskets by removing screws attaching adapter to wing and fuel cell.
On aircraft equipped with long-range cells, remove
cover plate and gaskets, and remove nylon vent tube
from inside cell.
h. Working through filler neck opening, loosen snap
fasteners. Tilt snap fasteners slightly when pulling
cell free, to prevent tearing rubber.
i. Collapse and carefully fold cell for removal,
then work cell out of fuel bay through filler opening
In upper wing surface. Use care when removing to
prevent damage to cell.
j. Unfold cell and remove fittings, snap fasteners
and fuel sump drain adapter.
13-12.

f. Place cell in compartment, develop it out to full
size and attach fasteners, then reverse procedures
outlined in the preceding paragraph for installation.
Install all new gaskets when installing cell.
g. On aircraft equipped with long-range cells, install nylon vent tube inside cell, inserting tube
through four hangers in top of cell. If a replacement
cell is being installed, use nylon vent tube removed
from old cell or order tube from applicable Parts
Catalog.
h. When tightening screw-type clamps, apply a
maximum of 20 pound-inches torque to clamp screws.
No oil is to be applied to fittings prior to installation.
i. When installing filler adapter, cover plate and
fuel quantity transmitter to the wing and fuel cell,
tighten attaching screws evenly. The sealing or compression surfaces must be assembled when absolutely
dry (NO SEALING PASTE IS TO BE USED).
j. After installation has been completed, cell should
be inspected for final fit within compartment, making
certain that cell is extended out to the structure and
no corners are folded in.
k. The final inspection, prior to closing the cell,
should be a close check to ensure that cell is free of
foreign matter such as lint, dust, oil or any installation equipment. If a cell is not thoroughly clean, it
should be cleaned with a lint-free cloth, soaked in
water, alcohol or kerosene. NO OTHER SOLVENT
SHALL BE USED.

FUEL CELL REPAIR.
NOTE

For fuel cell repair information, refer to
Cessna Service News Letter dated August
28, 1970. For minor repair, a fuel cell
repair kit is available from Goodyear,
complete with required materials and instructions.

NOTE
Throughout the aircraft fuel system, from
the cells to the engine-driven fuel pump,
use NS-40 (RAS-4) (Snap-On Tools Corp.,
Kenosha, Wisconsin), MIL-T-5544 (Thread
Compound, Antiseize, Graphite-Petrolatum)
or equivalent compound as a thread lubricant or to seal a leaking connection. Apply
sparingly to male fittings only, omitting

the first two threads.

Always ensure that

13-13.

Deleted.

13-14.

Deleted.

13-15.

Deleted.

13-20.

FUEL QUANTITY TRANSMITTERS.

13-16.

Deleted.

13-21.

DESCRIPTION.

a compound, the residue from a previously

used compound, or any other foreign material cannot enter the system.

Two fuel quantity indicators.

located in a cluster on the instrument panel are act13-17.

Deleted.

13-18.

Deleted.

13-19.

FUEL CELL INSTALLATION.

13-22.

a. Cell compartment must be thoroughly cleaned of
all filings, trimmings, loose washers, bolts, nuts,
etc.
b. All sharp edges of cell compartment must be
rounded off and protective tape applied over any other
sharp edges and protruding rivets.
c. Inspect cell compartment just prior to installa13-12

uated individually by an electric fuel quantity transmitter installed in each fuel cell.

Change 1

REMOVAL AND INSTALLATION.

(Refer to

Section 16.)
13-23. REMOVAL AND INSTALLATION OF FUEL
RESERVOIR TANKS.
a. Remove front seats, carpeting, and access
plates as necessary for access to tank to be removed.
b. Disconnect fuel lines at the tank to be removed.
c. Remove four screws securing tank mounting

Hinge for vent valve (11) must be at top. Tube
for vent extends into fuel cell, then is offset
upward.

Detail

A

3-

10

2

-

Detail B

-

FUEL
SAMPLER CUP

/

paragraph 2-19)

A

-

DetailD

1

C

*
STANDARD CELL

/
13

16
1. Plug/Valve
2. Gasket
3. Adapter
4. Clamp
5. Fitting
6. Wing Skin

7.
8.
9.
10.
11.

Filler Cap
Vent Line
Grommet
Hose
Vent Valve

Figure 13-5.

12.
13.
14.
15.
16.

Ground Strap
Fuel Quantity
Transmitter
Hanger (Typ)
Strainer
Protector

12
Detail C
FUEL QUANTITY TRANSMITTER
INSTALLATION AND GROUNDING

Fuel Cell Installation (Sheet 1 of 2)
Change 2

13-13

Hinge for vent valve (12) must be at top. Tube
for vent extends into fuel cell, then is offset
upward.
9

DetailB

A

Detail

3

10

I

I

2

B

FUEL
SAMPLER CUP
(Refer to

A paragraph 2-19)

2

17

v

2 14
C

Detail

D
LONG - RANGE CELL

15

1.
2.
3.
4.
5.
6.

Plug/Valve
Gasket
Adapter
Clamp
Fitting
Wing Skin

14

7.
8.
9.
10.
11.
12.

Cover Plate
Filler Cap
Vent Line
Grommet
Hose
Vent Valve

Figure 13-5.
13-14

Change 2

13.
14.
15.
16.
17.

Ground Strap
Fuel Quantity
Transmitter
Strainer
Protecter
Hanger (Typ)

Detail C
FUEL QUANTITY TRANSMITTER
INSTALLATION AND GROUNDING

Fuel Cell Installation (Sheet 2 of 2)

legs
d.
e.
voir

to fuselage structure.
Lift out the tank.
Reverse the preceding steps to install a resertank.

13-24. REMOVAL AND INSTALLATION OF FUEL
SELECTOR VALVE.
a. Drain all fuel from wing tanks at fuel tank sump
drain plugs. With valve turned to LEFT TANK, drain
left fuel lines at selector valve; with valve turned to
RIGHT TANK, drain right fuel lines.
b. Remove control pedestal cover. (Refer to section 11 for procedures.)
c. Remove access hole covers in floorboard and
fuselage skin in area of fuel selector valve,
d. Disconnect all fuel lines from selector valve.
e. Disconnect square shaft from valve by removing
attached roll pin.
f. Remove bolts or screws attaching valve to support bracket and remove valve.
g. Install valve by reversing this procedure.
13-25. FUEL SELECTOR VALVE REPAIR. (See
figure 13-6. ) The fuel selector valve may be repaired by disassembly, replacement of defective
parts, and reassembly as follows:
a. Mark sump plate (23) and body (1) to ensure
correct reassembly, then remove sump plate (23)
and O-ring (22) after removing four screws.
b. Drive out roll pin (5) securing yoke (6) to shaft.
As yoke is lifted off, balls (8) and springs (7) are
free. Retain them.
c. Lift off washer (9).
d. Mark cover (4) and body to assure later alignment ot parts and remove screws (3).
e. With fine emery paper. sand off any burrs or
sharp edges on shaft (21). Apply petrolatum to
shaft as a lubricant. then work cover off shaft.
f. Drive back roll pin (13) and remove rotor (12).
Teflon seal (14), O-rings (15), washers (16), and
springs (17) are now free to be removed. Check all
parts carefully to locate any defects.
g. Remove burrs or sharp edges on shaft, lubricate and slide it down. out of body (1). Remove
teflon seals (20) and O-rings (19).
h. Remove O-ring (18) within body and O-ring (10)
within cover.
i. Replace all O-rings, lap or replace teflon seals,

and lubricate O-rings before installation.
{CAUTION
Install all parts in the relative position depicted in figure 13-6, otherwise the valve
will not operate correctly.
j. Install O-ring (18) in body shaft hole. Install
O-rings (19) and teflon seals (20), then slide shaft
and rotor into place. Position rotor in exact relalive position shown in figure 13-6, then install Oring (22) and sump plate (23)
k. Install . 169" diameter pins in body ports, then
slide springs (17), washers (16), O-rings (15) and
teflon seals over pins. Slide rotor (12) over shaft.
Remove .169" dia. pins and, readjusting rotor vs.
shaft position as necessary, tap roll pin (13) into
place, letting it protrude on the side depicted.
NOTE
This roll pin serves also as a stop, limiting
valve shaft travel.
1. Install O-ring (10) in cover, lubricate shaft (21)
with petrolatum, install large O-ring (11), and slide
cover down into place.

ICAUTION|
Make sure cover is installed in relative position illustrated. A lug on the cover protrudes
to serve as a stop detent and if the cover is
not installed correctly, the valve will not operate correctly.
m. Install brass washer (9) and yoke (6). Note the
position of the small hole in the squared, upper portion of the yoke. If this is reversed, the valve linkage will not attach properly.
13-26. AUXILIARY ELECTRIC FUEL PUMP. On
aircraft Serials U20601619 thru U20601632 and aircraft prior to Serial U20601605, the auxiliary electric fuel pump is mounted on either the left side or
right side of the firewall. On aircraft Serials U20601605 thru U20601618 and beginning with U20601633,
the auxiliary electric fuel pump is located under the
floorboard on the right side of cabin, immediately

SHOP NOTES:

Change 1

13-15

f-



Detail A

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

12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.

Valve Body
Lockwasher
Screw
Cover
Roll Pin
Yoke
Spring
Ball
Brass Washer
O-Ring
O-Ring

Rotor
Roll Pin
Seal
O-Rlng
Washer
Spring
O-Ring
O-Ring
Seal
Rotor
O-Ring

24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.

Screw
Washer
Spring
Cap
Screw
Washer
Handle
Placard
Selector Shaft
Grommet
Selector Valve

23. Sump Plate

Figure 13-6.

Fuel Selector Valve Assembly

forward of the copilot seat. An integral bypass and
check valve permits fuel flow through the pump even
when the pump is inoperative, but prevents reverse
flow. A separate overboard drain line from the pump
prevents entry of fuel into the electric motor, in the
event of pump internal leakage.
13-27.
13-16

REMOVAL AND INSTALLATION.

a.

Firewall mounted:
. Place fuel selector in OFF position.
2. Remove top half of cowl for access to pump.
3. Disconnect all fuel lines and electrical connections from pump.
4. Loosen clamps securing pump and lift pump
out.
5. Reverse preceding steps for installation.

SAFETY WIRE HOLE

NOTE
Torque nut (15) to 25-30 lb in.
SAFETY WIRE HOLE

1. Spring
2. Washer
3. Plunger
4. Top
5. Drain Control

6. Plate
7. O-Ring
8. Gasket
9. Filter
10. Retainer Ring

Figure 13-7.
Floor mounted:
1. Place fuel selector in OFF position.
2. Peel back carpet and remove access plate in
floorboard immediately forward of copilot seat.
3. Disconnect all fuel lines and electrical connections from pump.

11.

Standpipe
12. O-Ring
13. Bowl
14. O-Ring
15. Nut

Fuel Strainer

b.

4.

Loosen clamps securing pump and lift pump

5.

Reverse preceding steps for installation.

out.
13-28. ELECTRIC FUEL PUMP CIRCUITS. The
electric fuel pump circuit is operated by a split
13-17

rocker-type switch. The low side of the switch is
connected through the "START" position of the ignition switch so that the fuel pump will operate only
while the ignition switch is in the "START" position
and the low side of the fuel pump switch is turned on.
When the ignition key is released, the pump will stop.
The high side of the fuel pump switch will operate the
pump regardless of ignition switch position. A throttie shaft operated microswitch adds a resistance to
the high circuit to slow down the pump when the throttie is retarded to prevent an excessively rich mixture
as throttle is retarded while the electric pump is
operating in the high position. Refer to the following paragraph for rigging of the microswitch.
12-28A. DESCRIPTION. Thru Serial U20602199,
the electric auxiliary fuel pump, which supplies fuel
flow for starting and for engine operation if the
engine-driven fuel pump should fail, is controlled by
the auxiliary fuel pump switch, mounted on the instrument panel. The switch is a split-rocker type; the
right half positions are "HI," "LO" and off and the
left half positions are "MAX HI" and off. The right
half of the switch incorporates an intermediate "LO"
position used for normal starting, and a "HI" position
(when the top of the switch is fully depressed) for
vapor purging during hot engine starts. Maximum
fuel flow is produced when the left half of the switch
is held in the spring-loaded "MAX HI" position. In
the "MAX HI" position, an interlock within the switch
automatically trips the right half of the switch to its
"HI" position. When the spring-loaded left half of the
switch is released, the right half will remain in the
"HI" position until manually returned to the off position. With the right half of the switch in the "LO"
position, and the starter button depressed, the auxiliary fuel pump will operate at a low flow rate (providing proper fuel mixture for starting) as the engine
is being turned over with the starter.
NOTE
The auxiliary fuel pump will not operate in
the "LO" position until the starter button is
depressed.
With the right half of the switch in the "HI" position,
the pump operates at one of the two flow rates that
are dependent upon the setting of the throttle. With
the throttle open to a cruise setting, the pump is
operating at a high capacity to supply sufficient fuel
to maintain flight. When the throttle is moved toward
the closed position (as during letdown, landing and
taxiing), the fuel pump flow rate is automatically
reduced, preventing an excessively rich mixture
during these periods of reduced engine speed. When
the engine-driven fuel pump is functioning and the
auxiliary fuel pump is functioning and the auxiliary
fuel pump is turned on "HI", a fuel/air ratio considerably richer than the best power is produced unless the
mixture is leaned. If the auxiliary fuel pump switch
is accidently placed on "HI" (with master switch on)
with the engine stopped and the mixture rich, the
intake manifold will be flooded.

13-18

Change 3

12-28B. DESCRIPTION. Beginning with U20602200,
the yellow right half of the switch is labeled "START",
and its upper "ON" position is used for normal starting, minor vapor purging and continued engine operation in the event of an engine-driven pump failure.
With the right half of the switch in the "ON" position.
the pump operates at one of two flow rates that are
dependent upon the setting of the throttle. With the
throttle open to a cruise setting, the pump operates
high enough capacity to supply sufficient fuel flow to
maintain flight with an inoperative engine-driven fuel
pump. When the throttle is moved toward the closed
position (as during letdown, landing and taxiing), the
fuel pump flow rate is automatically reduced, preventing an excessively rich mixture during these
periods of reduced engine speed.
NOTE
If the engine-driven fuel pump is functioning
and the auxiliary fuel pump switch is placed
in the "ON" position, a fuel/air ratio considerably richer than best power is produced
unless the mixture is leaned. Therefore,
this switch should be turned off during takeoff.
CAUTION
If the auxiliary fuel pump switch is accidently
placed in the "ON" position with the master
switch on and the engine stopped, the intake
manifolds will be flooded.
The red left half of the switch is labeled "EMERG",
and its upper "HI" position is used in the event of an
engine-driven fuel pump failure during take-off or
high power operation. The "HI" position may also be
used for extreme vapor purging. Maximum fuel flow
is produced when the left half of the switch is held in
the spring-loaded "HI" position. In this position, an
interlock within the switch automatically trips the
right half of the switch to the "ON" position. When
the spring-loaded left half of the switch is released,
the right half will remain in the "ON" position until
manually returned to the "OFF" position.
13-29. RIGGING THROTTLE MICROSWITCH.
(Refer to figure 13-8. ) The aircraft is equipped with
a throttle-operated microswitch which slows down the
electric fuel pump whenever the throttle is retarded
while the electric pump is being used. The electric
fuel pump microswitch should slow down the pump as
the throttle is retarded to approximately 19 inches of
mercury manifold pressure (sea level aircraft) and
23 inches of mercury manifold pressure (turbocharged
aircraft).
NOTE
These settings must be established during
ground run-up only. These values will not
apply in flight.
a. Start engine and set throttle to obtain 19 inches
of mercury manifold pressure (sea level aircraft) or
23 inches of mercury manifold pressure (turbocharged

5

fuel pump rocker switch 'ON. "
d. Advance throttle to full open position.
e. Check metered fuel pressure/flow on ship's gage
for a flow of 88-96 pounds/hour (14. 7-16.0 gallons/
hour).
f. Adjust number one resistor (6) if required.
g. Retard throttle slowly from the full "OPEN" position
until the speed of the fuel pump can be audibly
detected to change due to microswitch activation.
h. Wait momentarily for the fuel flow gage to re-

spond.

i. The metered fuel pressure/flow on the ship's
gage should read on the low end red line or approximately one red line width above.
j. Adjust number two resistor (5) if required.
13-31. MAXIMUM HIGH BOOST CHECK. To verify
high position function, momentarily depress springloaded rocker and verify a noticeable increase in indicated fuel flow on the fuel flow gage.
1.
2.
3.
4.
5.

Throttle Shaft Lever Cam
Airbox Bracket
Switch Actuator
Microswitch
Mounting Screw

Figure 13-8. Rigging

Throttle Microswitch

aircraft)
b. Mark position of throttle control at instrument
panel and shut down engine.
at the engine throttle shaft
c. Adjust microswitch
lever as required to cause electric fuel pump to slow
down as the throttle is retarded to the marked position.
c. With mixture control in "IDLE CUT-OFF," electric fuel pump switch in "HI." and master switch in
"ON" position, listen for change in sound of electric
fuel pump as the throttle is retard to the marked po- top
sition.
13-20.

FUEL FLOW TEST.

(Refer to figure 13-9.)

NOTE
These tests are to be conducted with the
engine stopped and external power supplied
to the aircraft bus.
a. Apply 13.75 VDC .25V (27. 75 VDC ± . 25V) to
aircraft bus.
b. Set mixture control at "FULL RICH. "
c. Turn master switch "ON," and yellow auxiliary

13-32. FUEL STRAINER. The fuel strainer is located in the nose wheel well. Access to the strainer
is gained by removing fairings aft of the nose gear.
The fuel strainer drain control is located adjacent to
the oil dipstick. Access to the drain control is gained
through the oil dipstick cowling door.
13-33. FUEL STRAINER DISASSFMBLY. (Refer to
figure 13-7.) To disassemble and assemble the
strainer, proceed as follows:
off fuel selector valve.
a. Turn
b. Disconnect strainer drain tube and remove
safety wire, nut. and washer at bottom of filter
bowl and remove bowl.
c. Carefully unscrew standpipe and remove.
d. Remove filter screen and gasket. Wash filter
screen and bowl in solvent (Federal Specification
P-S-661. or equivalent) and dry with compressed
air.
e. Using a new gasket between filter screen and
assembly, install screen and standpipe. Tighten
standpipe only finger tight.
f. Using all new O-rings, install bowl. Note that
step-washer at bottom of bowl is installed so that
step seats against O-ring. Connect strainer drain
tube.
g. Turn on fuel selector valve, close strainer
drain, and check for leaks. Check for proper
operation.
h. Safety wire bottom nut to top assembly. Wire
must have right hand wrap. at least 45 degrees.
13-34. ELECTRIC FUEL QUANTITY INDICATORS.
AND TRANSMITTERS. Refer to Section 16 for description. removal. installation and calibration

SHOP NOTES:

Change 3

13-19

A-A
3 MOR-20-2 resistors

1

P206-0567 thru U206-061\

A

12 VOLT SYSTEM

-

* 2 AMOR-20-1.5 resistors
LOOKING AFT AT FIREWALL (LEFT-HAND SIDE)
12 VOLT SYSTEM BEGINNING WITH P206-0161 & U206-0438

*

A-A

Adjust AMOR20-1.5 to 0.5 * .05 prior to installation.
Readjust resistors as required to comply with re-

quirements as outlined in paragraph 13-30.

2 MZ-0020-031AV
resistors
P20600618 thru P20600647
U206-1438 thru U206-02199
12 VOLT SYSTEM

1 MOR-20-2 resistor
THRU P206-0566 & U206-1284

2 MOR-20-2 resistors

2
P206-0567 thru P20600647
U206-1285 thru
--- U206-02199

B-B
AR-25-20 resistor

U206-1573 thru U206-02199
B

10
_

* 2 AMOR-20-10 resistors

24 VOLT SYSTEM
Position slide on this resistor for maxi-

LOOKING AFT AT FIREWALL (LEFT-HAND SIDE) mum resistance (all the way to the end
opposite QD8 wire) (Refer to Section 20.
24 VOLT SYSTEM BEGINNING WITH U206-02200
1270625, page 7.1.2.)
6. High Boost Resistor (#1)
1. Fuse Holder
2. Battery Box
7. Adjustable Lug
* Adjust AMOR-20-10 to 6.2 * .03 ohms prior to
3. Battery Contactor
8. Jumper
installation. Readjust resistors as required to
comply with requirements as outlined in para9. Adjustable Lug
4. Diode
5. Low Boost Resistor (#2) 10. Bracket
graph 13-30.
Figure 13-9.

13-20

Change 3

Fuel Pump Resistors

SECTION 14
PROPELLERS AND PROPELLER GOVERNORS

TABLE OF CONTENTS
PROPELLERS . ...........
Description . . . . . . . ..
...............
Repair
.
Trouble Shooting ............
Removal
.14-3
Installation
...
......
PROPELLER GOVERNORS .........
Description
.............

14-1.

Page
.

. 14-1
14-1
14-1
14-2
14-3
14-3
14-3

PROPELLERS.

14-2. DESCRIPTION. The aircraft is equipped
with an all-metal, constant-speed, governor-regulated propeller. The constant-speed propeller is
single- acting, in which engine oil pressure, boosted
and regulated by the governor is used to obtain the
correct blade pitch for the engine load. Engine lubricating oil is supplied to the power piston in the propeller hub through the crankshaft. The amount and
pressure of the oil supplied is controlled by the enginedriven governor. Increasing engine speed will cause
oil to be admitted to the piston, thereby increasing
the blade pitch. Conversely, decreasing engine speed

...
.... .
14-8
Trouble Shooting .....
Removal . . . . . ..
. . . . ... .
14-8
14-8
Control Arm and Bearing Assembly.
Removal and Installation. .....
14-8
Installation .......
14-10
14-10
High-RPM Stop Adjustment .......
Rigging Propeller Governor Control . . 14-10

will result in oil leaving the piston, thus decreasing
the blade pitch.
14-3. REPAIR. Metal propeller repair first involves
evaluating the damage and determining whether the
repair will be a major or minor one. Federal Aviation Regulations, Part 43 (FAR 43), and Federal
Aviation Agency, Advisory Circular No. 43. 13 (FAA
AC No. 43. 13), define major and minor repairs, alterations and who may accomplish them. When making repairs or alterations to a propeller FAR 43,
FAA AC No. 43.13 and the propeller manufacturer's
instructions must be observed.

Change 1

14-1

14-4.

TROUBLE SHOOTING.
TROUBLE

FAILURE TO CHANGE PITCH.

PROBABLE CAUSE

REMEDY

Governor control disconnected or
broken.

Check visually.
place control.

Governor not correct for
propeller. (Sensing wrong.)

Check that correct governor is
installed. Replace governor.

Defective governor.

Refer to paragraph 14-9.

Defective pitch changing mechanism
inside propeller or excessive propeller blade friction.

Propeller repair or replacement
is required.

Improper rigging of governor
control.

Check that governor control arm
and control have full travel. Rig
control and arm as required.

Defective governor.

Refer to paragraph 14-9.

SLUGGISH RESPONSE TO
PROPELLER CONTROL.

Excessive friction in pitch
changing mechanism inside
propeller or excessive blade
friction.

Propeller repair or replacement
is required.

STATIC RPM TOO HIGH OR
TOO LOW.

Improper propeller governor
adjustments.

Perform static RPM check
Refer to section 12 and 12A
for procedures.

ENGINE SPEED WILL NOT
STABILIZE.

Sludge in governor.

Refer to paragraph 14-9.

Air trapped in propeller
actuating cylinder.

Trapped air should be purged
by exercising the propeller
several times prior to take-off
after propeller has been reinstalled or has been idle for an
extended period.

Excessive friction in pitch
changing mechanism inside
propeller or excessive blade
friction.

Propeller repair or replacement
is required.

Defective governor.

Refer to paragraph 14-9.

FAILURE TO CHANGE PITCH
FULLY.

SHOP NOTES:

14-2

Change 1

Connect or re-

14-4.

TROUBLE SHOOTING (Cont.)
TROUBLE

PROBABLE CAUSE

OIL LEAKAGE AT PROPELLER MOUNTING FLANGE.

OIL LEAKAGE AT ANY
OTHER PLACE.

REMEDY

Damaged O-ring and seal between
engine crankshaft flange and
propeller.

Check visually. Remove propeller
and install O-ring seal.

Foreign material between
engine crankshaft flange and
propeller mating surfaces or
mounting nuts not tight.

Remove propeller and clean
mating surfaces; install new
O-ring and tighten mounting
nuts evenly to torque value
in figure 14-1.

Defective seals, gaskets,
threads, etc., or incorrect
assembly.

Propeller repair or replacement
is required.

14-5. REMOVAL. Refer to figure 14-1.
a. Remove spinner attaching screws (2) and remove
spinner (1). spinner support (3) and spacers (4). Retain spacers (4).
b. Remove cowling as required for access to
mounting nuts (9).
c. Loosen all mounting nuts (9) approximately
1/4 inch and pull propeller (15) forward until stopped
by nuts.
NOTE
As the propeller (15) is separated from the
engine crankshaft flange, oil will drain
from the propeller and engine cavities.
d. Remove all propeller mounting nuts (9) and
pull propeller forward to remove from engine crankshaft (12).
e. If desired, the spinner bulkhead (11) can be removed by removing screws (10) attaching lugs (8) or
bolts (19) attaching bulkhead (11) to propeller.
14-6. INSTALLATION.
a. If the spinner bulkhead (11) was removed, position bulkhead so the propeller blades will emerge
from the spinner (1) with ample clearance and install spinner bulkhead attaching lugs and screws,
or bolts (19) and nuts attaching spinner bulkhead
to propeller.

d. Align propeller mounting studs and dowel pins
with proper holes in engine crankshaft flange and
slide propeller carefully over crankshaft pilot until
mating surfaces of propeller and crankshaft flange
are approximately 1/4 inch apart.
e. Install propeller attaching washers and nuts (9)
and work propeller aft as far as possible, then
tighten nuts evenly and torque to 660-780 lb-in.
f. Install any spacers (4) used between spinner
support and propeller cylinder, then install spinner
support and spinner. The spacers are used as required to cause a snug fit between the spinner (1)
and the spinner support (3).
14-7.

PROPELLER GOVERNORS.

14-8. DESCRIPTION. The propeller governor is a
single-acting, centrifugal type, which boosts oil pressure from the engine and directs it to the propeller
where the oil is used to increase blade pitch. A
single-acting governor uses oil pressure to effect a
pitch change in one direction only; a pitch change in
the opposite direction results from a combination of
centrifugal twisting moment of rotating blades and
compressed springs. Oil pressure is boosted in the
governor by a gear type oil pump. A pilot valve, fly
weight and speeder spring act together to open and
close governor oil passages as required to maintain
a constant engine speed.
NOTE

CAUTION
Avoid scraping metal from bore of spinner
bulkhead and wedging scrapings between
engine flange and propeller. Trim the inside diameter of the bulkhead as necessary
when installing a new spinner bulkhead.
b. Clean propeller hub cavity and mating surfaces
of propeller and crankshaft.
c. Lightly lubricate a new O-ring (13) and the crankshaft pilot with clean engine oil and install the O-ring
in the propeller hub.

Outward physical appearance of specific
governors is the same, but internal parts
determine whether it uses oil pressure to
increase or decrease blade pitch. The
propellers used on these aircraft require
governors which "sense" in a certain manner. "Sensing" is determined by the type
pilot valve installed inside the governor.
Since the basic governor may be set to
"sense" oppositely, it is important to
ascertain that the governor is correct for
the propeller being used.
14-3

NOTE
Use spacers (4) as required to
ensure a snug fit between spinner
(1) and spinner support (3).
Torque propeller mounting nuts
(9) to 660 - 780 lb-in.

Spinner
1.
2.
3.
4.
5.

Stud

6. Screw
7.
8.
9.

14-4

Change 1

Lug
Mounting Nut

10.

Screw

11.
12.
13.
14.
15.
16.
17.
18.
19.
20.

Spinner Bulkhead
Engine Crankshaft
O-Ring
Dowel Pin
Propeller
Tube
Safety Wire
Ring
Bolt
Washer

TWO-BLADED PROPELLER

Figure 14-1.

Screw
Spinner Support
Spacer
Cylinder

Propeller Installation (Sheet 1 of 4)

12

TWO-BLADED, EXTENDED HUB PROPELLER

Figure 14-1.

Propeller Installation (Sheet 2 of 4)

14-5

14-6

Change I

THREE-BLADED,
Figure 14-1.

EXTENDED HUB PROPELLER

Propeller Installation (Sheet 4 of 4)
14-7

1
2

USED ON TURBOCHARGED ENGINES

3

AND NON-TURBOCHARGED ENGINE
5Q
THE MODEL U206

/ON

2.
3.
4.

High RPM Stop Screw
Bearing Race
Control Arm

J

5. Nylon Bearing
6.
7.

/

l

I

AT'-V//

Rivet
Retainer

8. Screw
9. Governor Arm

Figure 14-2. Governor Control Arm and Bearing Assembly

14-9.

TROUBLE SHOOTING.

When trouble shoot-

ing the propeller-governor combination, it is recommended that a governor known to be in good condition
be installed to check whether the propeller or the
governor is at fault. Removal and replacement, rigging, high-speed stop adjustment, desludging and replacement of the governor mounting gasket are not
major repairs and may be accomplished in the field.
Repairs to propeller governors are classed as propeller major repairs in Federal Aviation Regulations,
which also define who may accomplish such repairs.
14-10. REMOVAL.
a. Remove cowling, nose cap and engine baffles
as required for access to governor.
b. Disconnect governor control from governor.
NOTE
Note EXACT position of all washers so that
washers may be installed in the same position on reinstallation.
c. Disconnect intake manifold balance tube at
front of engine and move as required for clearance.
d. Remove nuts and washers securing governor to
engine and pull governor from mounting studs.
e. Remove gasket from between governor and engine mounting pad.
14-11.

CONTROL ARM AND BEARING ASSEMBLY.

Refer to figure 14-2.

14-8

Change I

14-12.

REMOVAL AND INSTALLATION.

a. Using a scribe, make aligning index marks on
governor arm (9) and end of governor serrated shaft.
NOTE
The governor arm (9) must be installed on the
governor shaft in the same serration or the
governor speed will be changed approximately
200 rpm.
b. Remove safety wire from governor arm screw
and from screws attaching governor head to governor.
c. Remove screws (8) that pass through the nonnotched holes in the retainer (7).
d. Loosen, but do not remove, the four remaining
screws so that retainer (7) may be rotated.
e. Loosen screw in governor arm (9) so that arm
may be slipped toward end of serrated shaft.
f. Slip governor arm toward end of serrated shaft
and work retainer (7) and control arm (9) from governor (1).
NOTE
If governor arm (9) becomes disengaged from
serrated shaft, align index marks and install
arm on serrated shaft. The control arm
spring has approximately 1-1/2 turns preload.
g. Rotate and remove bearing race (3) from governor (1).

TYPE

A

1

USED ON NON-TURBOCHARGED
ENGINE ON THE MODEL P206

1. Propeller Governor
2. High-RPM Stop Screw
3. Governor Arm Extension
4. Nut
5. Control Rod End
6. Governor Control

TYPE

B

USED ON TURBOCHARGED ENGINES
AND NON-TURBOCHARGED ENGINE
ON THE MODEL U206

1.
2.
3.
4.
5.
6.

Propeller Governor
High-RPM Stop Screw
Arm and Bearing Assembly
Nut
Control Rod End
Governor Control

3

5

4

REFER TO FIGURE 14-2

Figure 14-3.

Governor and Control Adjustments
Change 1

14-9

h.

Reverse the preceding steps for reinstallation.

14-13. INSTALLATION.
a. Wipe governor and engine mounting pad clean.
b. Install a new gasket on the mounting studs. Install gasket with raised surface of the gasket screen
toward the governor.
c. Position governor on mounting studs, aligning
governor drive splines with splines in the engine and
install mounting nuts and washers. Do not force
spline engagement. Rotate engine crankshaft slightly
and splines will engage smoothly when properly
aligned.
d. Connect governor control to governor and rig
control as outlined in paragraph 14-15.
e. Connect intake manifold balance tube, if removed,
Ensure all clamps are tight.
f. Reinstall all items removed for access.
14-14. HIGH-RPM STOP ADJUSTMENT. Refer to
figure 14-3.
a. Remove engine cowling.
b. (TYPE B.) Disconnect cabin heater inlet air
duct from nose cap.
c. (TYPE A.) Remove plug button from left front
baffle.
d. Remove safety wire and loosen the high-speed
stop screw locknut.
e. Turn the stop screw IN to decrease maximum
rpm and OUT to increase maximum rpm. One full
turn of the stop screw causes a change of approximately 25 rpm.
f. Tighten stop screw locknut, safety wire stop
screw and make propeller control linkage adjustment
as necessary to maintain full travel.
g. Install cabin heater inlet air duct or plug button
and install cowling.
h. Test operate propeller and governor.
NOTE
It is possible for either the propeller low
pitch (high-rpm) stop or the governor highrpm stop to be the high-rpm limiting factor.
It is desirable for the governor stop to limit

the high-rpm at the maximum rated rpm for
a particular aircraft. Due to climatic conditions, field elevation, low-pitch blade angle
and other considerations, an engine may not
reach rated rpm on the ground. It may be
necessary to readjust the governor stop after
test flying to obtain maximum rated rpm when
airborne.
14-15. RIGGING PROPELLER GOVERNOR CONTROL.
a. Disconnect control end (5) from governor (1).
b. Place propeller control in cabin, full forward,
then pull it back approximately 1/8 inch and lock in
this position. This will allow "cushion" to assure
full contact with governor high-rpm stop screw.
c. Place governor arm against high-rpm stop
screw.
d. Loosen jam nuts and adjust control rod end
until attaching holes align while governor arm is
against high-rpm stop screw. Be sure to maintain
sufficient thread engagement of the control and rod
end. If necessary, shift control in the clamps to
achieve this.
e. Attach rod end to the governor. Be sure all
washers are installed correctly.
f. Operate the control to see that the governor arm
bottoms out against the low pitch stop and bottoms
out against or a maximum of . 12" from the high pitch
stop on the governor before reaching the end of control cable travel.
NOTE
Non-turbocharged engines on the Model P206
are equipped with an offset extension to the
governor arm. The offset extension has an
elongated slot to permit further adjustment.
The preceding steps may still be used as an
outline in the rigging procedure. The result
of rigging, in all cases, is full travel of the
governor arm (bottom out against both high
and low pitch stops) with some "cushion" at
both ends of control travel.
* Refer to the inspection chart in Section 2
for inspection and/or replacement interval for the propeller control.

SHOP NOTES:

14-10

Change 1

SECTION 15
UTILITY SYSTEMS

TABLE OF CONTENTS

Page

UTILITY SYSTEMS ............
............
Heating System
............
Description
Operation .............
.........
Trouble Shooting
Removal and Installation of
.....
.......
Components
. . .........
Defroster System
Description ......
.............
Operation
. .....
Trouble Shooting
Removal and Installation of
...........
Components
Ventilating System ...........
.
........
Description .
.
. . ..
..
. ..
. ..
Operation
Trouble Shooting ..........
Removal and Installation of
......
.........
Components
.
..
............
Oxygen System

15-1
15-1
-15-1
15-1
15-1

15-1.

UTILITY SYSTEMS.

15-2.

HEATING SYSTEM.

15-3
15-3
15-3
15-3
15-3
15-3
15-3
15-3
15-3
15-3
15-3
15-3

15-3. DESCRIPTION. On non-turbocharged aircraft. the heating system is comprised of the heat
exchange section of the left exhaust muffler, a heater valve. mounted on the left forward side if the
firewall, a duct across the aft side of the firewall,
a push-pull control up the instrument panel, and flexible ducts connecting the system. On aircraft with
turbocharged engines, the heating system consists of
an opening in the left side of the nose cap, an exhaust
shroud, a heater valve, mounted on the left forward
side of the firewall, to which is attached an adapter
and a tube extending downward and overboard. The
system also includes a duct across the aft side of the
firewall, a push-pull control on the instrument panel,
and flexible ducts connecting the system.
15-4. HEATER OPERATION. On airplanes with
non-turbocharged engines, ram air is ducted through
an engine baffle and the heat exchange section of the
left exhaust muffler, to the heater valve at the firewall. On aircraft with turbocharged engines, ram
air is ducted through an opening in the left side of the
nose cap, through an exhaust shroud, to the heater
valve at the firewall. On both models, heated air
flows from the heater valve into a duct across the aft
side of the firewall, where it is distributed into the
cabin. The heater valve, operated by a push-pull

............
Description
Maintenance Precautions .......
Replacement of Components
Oxygen Cylinder General
Information .....
Oxygen Cylinder Service
.
Requirements ......
Oxygen Cylinder Inspection
...
..
Requirements ..
Oxygen System Component
Service Requirements .15-10
Oxygen System Component
Inspection Requirements ...
Masks and Hose ....
...
Maintenance and Cleaning .
. . ...
System Purging ...
Functional Testing .....
System Leak Test .........
.........
System Charging

15-5
15-6
15-6
15-6
15-10
15-10

15-10
. 15-10
15-10
. 15-11
15-11
. 15-11
. 15-11

control marked "CABIN HEAT", located on the instrument panel, regulates the volume of heated air
entering the system. Pulling the heater control full
out supplies maximum flow, and pushing it in gradually decreases flow, shutting off flow completely
when the control is pushed full in.
15-5. TROUBLE SHOOTING. Most of the operational troubles in the heating system are caused by
sticking or binding air valves and their controls,
damaged air ducting, or defects in the exhaust muffler. In most cases, valves or controls can be freed
by proper lubrication. Damaged or broken parts
should be repaired or replaced. When checking controls, be sure valves respond freely to control movement, that they move in the correct direction, and
that they move through their full range of travel and
seal properly. Check that hose are properly secured
and replace hose that are burned, frayed or crushed.
If fumes are detected in the cabin, a very thorough
inspection of the exhaust muffler should be accomplished. Refer to the applicable paragraph in Section
12 for the non-turbocharged engine exhaust system
inspection, or for the turbocharged engine, refer to
Section 12A. Since any holes or cracks may permit
exhaust fumes to enter the cabin, replacement of defective parts is imperative because fumes constitute
an extreme danger. Seal any gaps in heater ducts
across the firewall with Pro-Seal #700 (Coast ProSeal Co., Los Angeles, California) compound, or
equivalent compound.
15-1

THRU P20600644 & U20601614
*BEGINNING
WITH U20601615

-....

3*

2A

12

"

.......

. . ... v
D'' :......
7

\,g

:::-.

..'

-........
I·
::.....

14\\f '

A

i»

DetailB

212
>

NON-TURBOCHARGED

20

\\

;'S

*NEOPRENE COATED ASBESTOS SEAL */
AND STAINLESS STEEL DOUBLER
BEGINNING WITH U20601637

1.
2.
3.
4.
5.
6.
7.
8.
9.

26

/

a

/2
31
29

/N

Retainer
Defroster Deflector
Cowl Deck
Defroster Outlet
Washer
Cotter Pin
Nut
Valve
Screw

17.
18.
19.
20.
21.
22.
23.
24.
25.

10.

Arm

26.

Valve Seat

AND TBE

11.
12.

Clamp Bolt
Shaft

27.
28.

Valve Extension
Reinforcement

USED ON
TURBO-

13.

Right Air Duct

29.

Shim

14.

Cabin Heat Control

30.

Spring

15.
16.

Defroster Control
DefrosterHose

31.
32.
33.

Block
Roll Pin
Deflector Plate

Left Air Duct
Tube
Adapter
Clamp
Hose
Shroud
Ram Air Intake
Valve Plate
Valve Body

;'<.21
i
21
*19

*

ADAPTER
* ADAPTER

CHARGED

_

3.

ENGINES

s
9

Detail C

Figure 15-1.

15-2

Change 3

Heating and Defrosting Systems (Sheet I of 2)

44

38

46

39

40

41

42

Detail

F

48
49
/51

-

52

54

Detail G

34.
35.
36.
37.
38.
39.
40.

Washer
Valve and Nozzle
Arm Assembly
Roll Pin
Shaft
Valve Assembly
Sta-Strap

Figure 15-1.

41.
42.
43.
44.
45.
46.
47.
48.

Cover
Plenum
Retainer
Outlet Cover
Spacer
Screen
Hose
Clamp

49.
50.
51.
52.
53.
54.
55.

ControlArm
Spring
Valve Plate
Valve Seat
Nutplate
Valve Extension
Valve Body

Heating and Defrosting Systems (Sheet 2 of 2)
Change 3

15-2A/(15-2B blank)

15-6. REMOVAL AND INSTALLATION OF COMPONENTS. Figure 15-1 may be used as a guide for
removal and installation of components of the heater
system. Cut replacement hose to length and install
in the original routing. Trim hose winding shorter
than the hose to allow hose clamps to be fitted. Defective heater valves should be repaired or replaced.
Check for proper operation of valves and their controls after installation or repair.
15-7.

DEFROSTER SYSTEM.

15-8. DESCRIPTION. The system is composed of
a duct across the aft side of the firewall, a defroster
outlet, mounted in the left side of the cowl deck immediately aft of the windshield, a defroster control
knob on the instrument panel, and flexible ducting
connecting the system.
15-9. DEFROSTER OPERATION. Air from the duct
across the aft side of the firewall flows through a
flexible duct to the defroster outlet. The defroster
control operates a damper in the outlet to regulate
the amount of air deflected across the inside surface
of the windshield. The temperature and volume of
this air is controlled by the settings of the cabin
heating system ontrol.
15-10. TROUBLE SHOOTING. Most of the operational troubles in the defrosting system are caused
by sticking or binding of the damper in the defroster
outlet or its control.
Since the defrosting system
depends on proper operation of the cabin heating system. refer to paaragraph 15-5 for trouble shooting the
heating and defrosting system.
15-11. REMOVAL AND INSTALLATION OF COMPONENTS. Figure 15-1 may be used as a guide for
removal and installation of components of the defrosting system. Cut replacement hose to length and
install in the original routing. Trim hose winding
shorter than the hose to allow hose clamps to be fitted.
A detective defroster outlet should be repaired
or replaced. Check
for proper operation of defroster
outlet and its control after installation or repair.
15-12.

VENTILATING SYSTEM.

15-13. DESCRIPTION. The system is comprised of
two airscoops mounted in the inboard leading edge of
each wing, an adjustable ventilator mounted on each
side of the cabin near the upper corners of the windshield, two plenum chambers mounted in the left and
right rear cabin wing root areas, two fresh airscoop
doors, one on each side of the fuselage, just forward
of the front seats, a control on the instrument panel
for each of these scoop doors and flexible ducting
connecting the system.
15-14. VENTILATING SYSTEM OPERATION. Air
received from scoops mounted in the inboard leading
edges of the wings is ducted to adjustable ventilators
mounted on each side of the cabin near the upper corners of the windshield. Rear seat ventilation is provided
by plenum chambers mounted in the left and
right rear cabin wing root areas. These plenum
chambers receive ram air from the airscoops in the

inboard leading edges of the wings. Each plenum
chamber is equipped with a valve which meters the
incoming cabin ventilation air. This provides a
chamber for the expansion of cabin air which greatly
reduces inlet air noise. Filters at the air inlets are
primarily noise reduction filters. Forward cabin
ventilation is provided by two fresh airscoop doors,
one on each side of the fuselage, just forward of the
front seats. The left scoop door is operated by a
control in the instrument panel marked "CABIN AIR. "
and the right scoop door is operated by a control in
the instrument panel marked "AUX CABIN AIR. "
Fresh air from the scoop doors is routed to the duct
across the aft side of the firewall, where it is distributed into the cabin. As long as the "CABIN
HEAT" control is pushed full in, no heated air can
enter the firewall duct; therefore, when the "CABIN
AIR" or "AUX CABIN AIR" controls are pulled out,
only fresh air from the scoops will flow through the
duct into the cabin. As the "CABIN HEAT" control
is gradually pulled out, more and more heated air
will blend with the fresh air from the scoops and be
distributed into the cabin. All of the controls may
be set at any position from full open to full closed.
15-15. TROUBLE SHOOTING. Most of the operational troubles in the ventilating system are caused
by sticking or binding of the lever in the inlet scoop
door or its control. The spring or plate in the plenum chambers could also bind or stick, requiring
repair or replacement of the plenum chamber.
Check the filter elements in the airscoops in the
leading edges of the wings for obstructions. The
elements may be removed and cleaned or replaced.
Since air passing through the filters is emitted into
the cabin, do not use a cleaning solution which would
contaminate cabin air. The filters may be removed
to increase air flow. However, their removal will
cause a slight increase in noise level.
15-16. REMOVAL AND INSTALLATION OF COMPONENTS. Figure 15-2 may be used as a guide for
removal and installation of components of the ventilating system. Cut replacement hose to length and
install in the original routing. Trim hose winding
shorter than the hose to allow hose clamps to be fitted. A defective plenum chamber should be repaired
or replaced. Check for proper operation of ventilating controls after installation or repair.
15-17.

OXYGEN SYSTEM.

WARNING
Under NO circumstances should the ON-OFF
control on the oxygen regulator be turned to
the "ON" position with the outlet (low pressure) ports open to atmosphere. Operation
of these units in this manner will induce
serious damage to the regulators and having
the following results:
1. Loss of outlet set pressure.
2. Loss of oxygen flow through the regulator which will result in inadequate oxygen being fed
through the aircraft system.
3. Internal leakage of oxygen through the
regulator.
15-3

SEE SHEET 2

4

DetA i

et

4il

K

·Detail
SA
:;-

1

o5
I

a

9

4
BEGINNING WITH U20601661

BEINNNGSTALLATION
WIT
AIRCRAFT SERIAL

1

*I~^
21,T

20,

' '

^'.

Detai

U.'3

' A

1.
2.
3.
4.
5.
6.
7.
8.

"'
T20U U206

5rU20602236

4. AirspeedIndicator

.1

12. Sump
Leit Sump)WSystems
TH-c
Insert 15.toBEGINNING
14.(Airspeed
6 Line
Figur
SERIAL
DetailB
AIRCRAFT
U20602236

Line (To Right Sump)
Altimeter
Vertical Speed Indicator
Airspeed Indicator
Line (To Pitot Tube)
Line (Airspeed to Left Sump)
Stringer
Nutplate

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

Spacer
Clamp
Screw
Sump
Line (Airspeed to Left Sump)
Insert 15.
Line (To Instruments)
Line (To Alternate Air)

Figure 16-2.

20. Static Port Mast Body
22. pitot Tube
BEGINNING WITH U20601661
THRU U20602235
17.
18.
19.
20.
21.
22.
23.
24.

Bracket
Line (To Sumps)
Valve
Static Port
Connector
pitot Tube Mast Body
Heater Element
Sta-Strap

Pitot-Static Systems
Change 1

16-3

NOTE
Do not overtighten screws (2) and do not
lubricate any parts.

6

Use spacers (6) as required for adequate
friction on ring assembly (4).

\

4

1

NOTE
Specific airspeed indicators, listed
by part number in applicable Parts
Catalogs, must be used in the true airspeed installation. Internal mechanism,
face plate, and calibration are different
from those of a standard instrument.

Figure 16-3.

1. Instrument Panel Cover

5. Instrument Panel

2. Mounting Screw

6. Spacer

3.Retainer
4. True Airspeed Ring

7. Airspeed Indicator
8. Nut

True Airspeed Indicator

NOTE
If panel is to be removed from aircraft,
remove control wheel.
d. To remove shock-mounted panel remove nuts
from shock mounts and pull panel straight back.
e. Reverse preceding steps for installation.
NOTE
A light coat of paraffin, beeswax or soap on
prongs of retainer clips will ease installation.
16-6. SHOCK MOUNTS. Service life of Instruments
is directly related to adequate shock-mounting of
panel. If removal of panel is necessary, check
mounts for deterioration and replace as necessary.
16-7.

INSTRUMENTS.

16-9. INSTALLATION. Generally, installation procedure is the reverse of removal procedure. Make
sure mounting screw nuts are tightened firmly, but
do not overtighten, particularly on instruments having plastic cases. The same rule generally applies
to connecting plumbing and wiring.
NOTE
All instruments (gages and indicators), requiring a thread seal or lubricant, shall be
installed using teflon tape on male fittings
only. This tape is available through Cessna
Service Parts Center.

(Refer to figure 16-1.)

16-8. REMOVAL. Most instruments are secured to
panel with screws inserted through panel face, under
decorative cover. To remove an instrument, remove
decorative cover, disconnect plumbing or wiring to
instrument concerned, remove retainer screws and
take instrument out from behind, or, in some cases
from front of instrument panel. Instrument clusters
are installed as units, secured by a screw on each
corner of cluster. Cluster must be removed from
panel to replace an individual gage. In all cases
when an Instrument is removed, lines or wires
disconnected from it should be protected. Cap open
lines and cover pressure connections on instrument
16-4

to prevent thread damage and entrance of foreign
matter. Wire terminals should be insulated or tied
up so they will not ground accidentally or shortcircuit on another terminal.

When replacing an electrical gage in an instrument
cluster assembly, avoid bending pointer or dial
plate. Distortion of dial or back plate could change
calibration of gages.
16-10. PITOT AND STATIC SYSTEMS.
figure 16-2.)

(Refer to

16-11. DESCRIPTION. The pitot system conveys
ram air pressure to the airspeed indicator. The
static system vents vertical speed indicator, altimeter and airspeed indicator to atmospheric pressure through plastic tubing connected to static ports.

A static line sump is installed at each source button
to collect condensation in static system. Beginning
with 1974 models a new smaller diameter static line
sump is installed and is located on the firewall. An
alternate static source may be installed and is used
only in emergencies. When used as a static source
on Aircraft Serials thru U20601632 the cabin air
becomes another source of static air and the external
source is not shut off unless totally obstructed. Beginning with Serial U20601633 the static source valve
is so connected to the system that when the control is
pulled on the external source is mechanically shut off
and the cabin air becomes the only source of static
air. When used as a static source, cabin pressure is
substituted for atmospheric pressure, causing instrument readings to vary from normal. Refer to Owner's
Manual for flight operation using alternate static
source pressure. A pitot tube heater and stall warning heater may be installed. The heating elements
are controlled by a switch at the instrument panel and
powered by the electrical system.
Proper maintenance of
16-12. MAINTENANCE.
pitot and static system is essential for proper operation of altimeter, vertical speed and airspeed indicators. Leaks, moisture and obstructions in pitot
system will result in false airspeed indications,
while static system malfunctions will affect readings
of all three instruments. Under instrument flight
conditions. these instrument errors could be hazardous. Cleanliness and security are the principal
rules for system maintenance. The pitot tube and
static ports MUST be kept clean and unobstructed.
16-13. STATIC PRESSURE SYSTEM INSPECTION
AND LEAKAGE TEST. The following procedure
outlines inspection and testing of static pressure
system, assuming altimeter has been tested and inspected in accordance with current Federal Aviation
Regulations.
a. Ensure static system is free from entrapped
moisture and restrictions.
b. Ensure no alterations or deformations of airframe surface have been made which would affect
the relationship between air pressure in static pressure system and true ambient static air pressure for
any flight configuration.
c. Seal off one static pressure source opening with
plastic tape. This MUST be an air-tight seal.
d. Close static pressure alternate source valve, if
installed.
e. Attach a source of suction to remaining static
pressure source opening. Figure 16-4 shows one
method of obtaining suction.
f. Slowly apply suction until altimeter indicates a
1000-foot increase in altitude.
CAUTION
When applying or releasing suction, do not
exceed range of vertical speed indicator or
airspeed indicator.
g. Cut off suction source to maintain a "closed"
system for one minute. Leakage shall not exceed

100 feet of altitude loss as indicated on altimeter.
h. If leakage rate is within tolerance, slowly release suction source, then remove tape used to
seal static source.
NOTE
If leakage rate exceeds maximum allowable.
first tighten all connections, then repeat
leakage test. If leakage rate still exceeds
maximum allowable, use following procedure.
i. Disconnect static pressure lines from airspeed
indicator and vertical speed indicator. Use suitable
fittings to connect lines together so altimeter is the
only instrument still connected into static pressure
system.
j. Repeat leakage test to check whether static pressure system or the removed instruments are cause of
leakage. If instruments are at fault, they must be
repaired by an "appropriately rated repair station"
or replaced. If static pressure system is at fault,
use following procedure to locate leakage.
k. Attach a source of positive pressure to static
source opening. Figure 16-4 shows one method of
obtaining positive pressure.
CAUTION
Do not apply positive pressure with airspeed
indicator or vertical speed indicator connected to static pressure system.
1. Slowly apply positive pressure until altimeter
indicates a 500-foot decrease in altitude and maintain this altimeter indication while checking for leaks.
Coat line connections, static pressure alternate
source valve and static source flange with solution of
mild soap and water, watching for bubbles to locate
leaks.
m. Tighten leaking connections. Repair or replace
parts found defective.
n. Reconnect airspeed and vertical speed indicators
into static pressure system and repeat leakage test
per steps "c" thru "h".
16-14. PITOT SYSTEM INSPECTION AND LEAKAGE
TEST. To check pitot system for leaks, fasten a
piece of rubber or plastic tubing over pitot tube, close
opposite end of tubing and slowly roll up tube until
airspeed indicator registers in cruise range. Secure tube and after a few minutes recheck airspeed
indicator. Any leakage will have reduced the pressure in system, resulting in a lower airspeed indication. Slowly unroll tubing before removing it, so
pressure is reduced gradually. Otherwise instrument may be damaged. If test reveals a leak in system, check all connections for tightness.
16-15. BLOWING OUT LINES. Although pitot system is designed to drain down to pitot tube opening,
condensation may collect at other points in system
and produce a partial obstruction. To clear line,
disconnect at airspeed indicator. Using low pressure air, blow from indicator end of line toward
pitot tube.
Change 1

16-5

PRESSURE

THICK-WALLED
SURGICAL HOSE

PRESSURE BLEED-OFF
SCREW (CLOSED)
AIR BULB
WITH CHECK
VALVES
CLAMP
CLAMP
THICK-WALLED
SURGICAL HOSE
CHECK VALVE
NOTE

CHECK VALVE

SUCTION

Air bulb with check valves may be obtained
locally from a surgical supply company.
This is the type used in measuring blood
pressure.

TO APPLY SUCTION:
1.

Squeeze air bulb to expel as much air as possible.

2.

Hold suction hose firmly against static pressure source opening.

3.

Slowly release air bulb to obtain desired suction, then pinch hose shut tightly to trap suction in
system.

4.

After leak test, release suction slowly by intermittently allowing a small amount of air to enter
static system. To do this, tilt end of suction hose away from opening, then immediately tilt it
back against opening. Wait until vertical speed indicator approaches zero, then repeat. Continue to admit this small amount of air intermittently until all suction is released, then remove
test equipment.

TO APPLY PRESSURE:

CAUTION
Do not apply positive pressure with airspeed indicator or vertical speed
indicator connected into static system.
1.

Hold pressure hose firmly against static pressure source opening.

2.

Slowly squeeze air bulb to apply desired pressure to static system. Desired pressure may be
maintained by repeatedly squeezing bulb to replace any air escaping through leaks.

3.

Release pressure by slowly opening pressure bleed-off screw, then remove test equipment.

Figure 16-4.

16-6

Static System Test Equipment

|CAUTIONI
Never blow through pitot or static lines toward
instruments.
Like pilot lines, static pressure lines must be kept
clear and connections tight. All models have static
source sumps which collect moisture and keep sys(em clear. However, when necessary, disconnect
static line at first instrument to which it is connected, then
then blow
blow line
line clear
clear with
with low-pressure
low-pressure air.
air.
ed,

On aircraft equipped with alternate static
source, use same procedure, opening
alternate static source valve momentarily
to clear line, then close valve and clear
remainder of system.
Check all static pressure line connections for tightness. If hoses or hose connections are used, check
for general condition and clamps for security. Replace hoses which have cracked, hardened or show
other signs of deterioration.
16-17.

16-16. REMOVAL AND INSTALLATION.
(Refer to figure 16-2.) To remove pitot mast
remove our mounting screws on side of
connector (21) and pull mast uut of connector tar
enough to disconnect pitot line (5). Electrical conneclions to heater assembly (if installed) may be
disconnected through wing access plate just inb.ard
of mast. Piot and static lines are removed in e
usual manner,
manner, after
alter removing
removing wing
wing access
access plates
plates.
usual
lower wing (airing strip and upholstery as requied.
iInstallation
ir of tubing will be simpler if a guide wiie
is drawn in as
tubing is removed from wing. The
tubing may be removed intact by drawing it out
through cabin and right door. When replacing components of pilot and static pressure systems. use
anti-seize compound sparingly on male threads on
both metal and plastic connections. Avoid excess
compound which might enter lines. Tighten connections firmly, but avoid overlighlening and distorting fittings. If twisting of plastic tubing is
encountered when tightening fittings, VV-P-236
encP etrolatume , may be applied sparingly between
tubing and fittings

TROUBLE SHOOTING--PITOT STATIC SYSTEM.
REMEDY

PROBABLE CAUSE

TROUBLE
LOW OR SLUGGISH AIRSPEED
INDICATION. (Normal altimeter
and vertical speed. )

Pitot tube obstructed, leak or
obstruction in pitot line.

Test pitot tube and line for leaks
or obstructions. Blow out tube
and line. repair or replace damaged line.

INCORRECT OR SLUGGISH
RESPONSE. (all three
instruments. )

Leaks or obstruction in static
line.

Test line for leaks and obstructions. Repair or replace line.
blow out obstructed line.

16-18. TRUE AIRSPEED INDICATOR. A true airspeed indicator may be installed. This indicator.
equipped with a conversion ring, may be rotated until
pressure altitude is aligned with outside air temperature, then airspeed indicated on instrument is read
as true airspeed on adlustable ring. Refer to ligure
16-3 for removal and installation. Upon installation,
before tightening mounting screws (2), calibrate the
instrument as follows: Rotate ring (4) until 120 mph
_~~

on adjustable ring aligns with 120 mph on indicator.
Holding this setting, move retainer (3) until 60 F
aligns with zero pressure altitude, then Lighten
mounting screws (2) and replace decorative cover.

~above

~~~~~~~~~~~~~ChangeSHOPNO

NOTE
On indicators graduated in knots, use 105
knots instead of 120 miles per hour in the

NOTES:
~~~SHOP
calibration

procedure.

ES:16-7

Change 3

16-7

16-19.

TROUBLE SHOOTING-AIRSPEED INDICATOR.
TROUBLE

HAND FAILS TO RESPOND.

INCORRECT INDICATION
OR HAND OSCILLATES.

PROBABLE CAUSE

Pitot
pressure connection not
properly connected to pressure line from pitot tube.

Test line and connection for leaks.
Repair or replace damaged line,
tighten connections.

Pitot or static lines clogged.

Check line for obstructions.
out lines.

Leak in pitot or static lines.

Test lines and connections for
leaks. Repair or replace damaged lines, tighten connections.

Defective mechanism or
leaking diaphragm.

Substitute known-good indicator
and check reading. Replace
instrument.

Leaking diaphragm.

(Refer to Paragraph 16-11)

HAND VIBRATES.

SHOP NOTES:

16-8

Change 1

REMEDY

Blow

Substitute known-good indicator
and check reading. Replace
instrument.

Alternate static source valve
open. THRU U20601596,
U20601619 THRU U20601632
AND THRU P20601587.

Check visually. Close for
normal operation.

Excessive vibration.

Check panel shock mounts. Replace defective shock mounts.

Excessive tubing vibration.

Check clamps and line connections
for security. Tighten clamps and
connections, replace tubing with
flexible hose.

16-20.

0

TROUBLE SHOOTING--ALTIMETER
TROUBLE

INSTRUMENT FAILS TO
OPERATE.

INCORRECT INDICATION.

HAND OSCILLATES.

16-21.

PROBABLE CAUSE

REMEDY

Static line plugged.

Check line for obstructions.
Blow out lines.

Defective mechanism.

Substitute known-good altimeter and check reading.
Replace instrument.

Hands not carefully set.

Reset hands with knob.

Leaking diaphragm.

Substitute known-good altimeter and check reading.
Replace instrument.

Pointers out of calibration.

Compare reading with knowngood altimeter. Replace
instrument.

Static pressure irregular.

Check lines for obstruction
or leaks. Blow out lines,
tighten connections.

Leak in airspeed or vertical
speed indicator installations.

Check other instruments and
system plumbing for leaks.
Blow out lines, tighten connections.

TROUBLE SHOOTING--VERTICAL SPEED INDICATOR.
TROUBLE

INSTRUMENT FAILS TO
OPERATE.

INCORRECT INDICATION.

POINTER OSCILLATES.

PROBABLE CAUSE

REMEDY

Static line plugged.

Check line for obstructions.
Blow out lines.

Static line broken.

Check line for damage, connections for security. Repair or replace damaged line,
tighten connections.

Partially plugged static line.

Check line for obstructions.
Blow out lines.

Ruptured diaphragm.

Substitute known-good indicator
and check reading. Replace
instrument.

Pointer off zero.

Reset pointer to zero.
pointer to zero.

Partially plugged static line.

Check line for obstructions.
Blow out lines.

Reset

16-9

16-21.

TROUBLE SHOOTING--VERTICAL SPEED INDICATOR.
TROUBLE

POINTER OSCILLATES.

PROBABLE CAUSE
(cont).

HAND VIBRATES.

16-22.

TUBE DOES NOT HEAT OR
CLEAR ICE.

Leak in static line.

Test lines and connections for
leaks. Repair or replace damaged lines, tighten connections.

Leak in instrument case.

Substitute known-good indicator
and check reading. Replace
instrument.

Excessive vibration.

Check shock mounts. Replace
defective shock mounts.

Defective diaphragm.

Substitute known-good indicator
and check for vibration. Replace instrument.

PROBABLE CAUSE

Turn switch "ON."

Blown fuse.

Check fuse.

Break in wiring.

Test for open circuit.
wiring.

Heating element burned out.

Check resistance of heating
element. Replace element.

VACUUM SYSTEM (Refer to Figure 16-5)

Replace fuse.
Repair

the system. A discharge tube is connected to the
pump to expell the air from the pump overboard. A
suction relief valve is used to control system pressure and is connected between the pump inlet and the
instruments. In the cabin, the vacuum line is routed
from the gyro instruments to the relief valve at the
firewall. A central air filtering system is utilized.
The reading of the suction gage indicates net difference in suction before and alter air passes through
a gyro. This differential pressure will gradually
decrease as the central air filter becomes dirty,
causing a lower reading on the suction gage.

TROUBLE SHOOTING--VACUUM SYSTEM -- THRU U20601956 (WET SYSTEM)
TROUBLE

HIGH SUCTION GAGE READINGS.

16-10

REMEDY

Switch turned "OFF."

16-24. DESCRIPTION. Through Aircraft Serial
U20601956 suction to operate the gyros is provided
by an engine-driven vacuum pump, gear-driven
through a spline-type coupling. The vacuum pump
discharge air passes through an oil separator, where
the oil, which passes through the pump for lubrication, is returned to the engine and the air is expelled
overboard. Beginning with Aircraft Serial U20601957
a dry vacuum system is installed. This system utilizes a sealed bearing, engine-driven vacuum pump,
which eliminates the oil separation components from
16-25.

REMEDY

TROUBLE SHOOTING--PITOT TUBE HEATER.

TROUBLE

16-23.

(Cont)

Change 1

PROBABLE CAUSE
Gyros function normally-relief
valve screen clogged, relief
valve malfunction.

REMEDY
Check screen, than valve.

Com-

pare gage readings with new gage.
Clean screen, reset valve.
place gage.

Re-

0

NOTE
Refer
relief valve
forparagraph
16-30 to
adjustment.

1. Oil Separator
2. Vent
3. Bracket
4. Oil Return (To Engine)
5. Vacuum Pump

7. Filter Element
8. Wing Nut
9. Suction Hose
10. Suction Gage
11. Directional Gyro

6. Bracket

12. Gyro Horizon

Figure 16-5.

13. Relief Valve
14. Vacuum Adjust
15. Tube Locator
16. Firewall
17. O-Ring

18.

Fitting

19.

Cross Assembly

Vacuum System (Sheet 1 of 3) Wet System
Change 3

16-11

NOTE
Detail
THRU U20603020

Refer to paragraph
16-30 for relief valve
adjustment.

2

7

18
16

*

Detail C

2*

BEGINNING WITH
AIRCRAFT SERIAL
U20601957

5
*FOR TU206 MODELS, VENT
TUBE IS POSITIONED AS SHOWN
Figure 18-5.
16-12

Change 3

(DRY SYSTEM)

Vacuum System (Sheet 2 of 3) Dry System

D

Detail D

BEGINNING WITH U20603021

Figure 16-5.

Vacuum System (Sheet 3 of 3) Dry System
Change 3

16-12A

16-25.

TROUBLE SHOOTING--VACUUM SYSTEM--THRU U20601956 (WET SYSTEM) (cont)
TROUBLE

NORMAL SUCTION GAGE
READING, SLUGGISH OR

PROBABLE CAUSE

REMEDY

Instrument air filters clogged.

Clean or replace filter as
necessary.

Leaks or restriction between
instruments and relief valve.
relief valve out of adjustment,
defective pump, restriction
in oil separator or pump
discharge line.

Check lines for leaks. disconnect
and test pump. Repair or replace
lines, adjust or replace relief
valve. repair or replace pump.
clean oil separator.

Central air filter dirty.

Clean or replace filter as
necessary.

SUCTION GAGE FLUCTUATES.

Defective gage or sticking relief
valve.

Check suction with test gage.
Replace gage. Clean sticking
valve with Stoddard solvent.
Blow dry and test. If valve
sticks after cleaning, replace
valve.

OIL COMES OVER IN PUMP
DISCHARGE LINE.

Oil seperator clogged, oil return
line obstructed, excessive oil flow
through pump.

Check oil seperator, return line.
Check that pump oil return rate
does not exceed 120 cc/hour
(approx. 8 drops/minute), at 50
psi oil pressure. Clean oil separator is Stoddard solvent, blow
dry. Blow out lines. If pump oil
consumption is excessive, replace oil metering collar and pin
in pump.

ERRATIC GYRO RESPONSE.

LOW SUCTION GAGE
READINGS.

16-25A.

TROUBLE SHOOTING--VACUUM SYSTEM--BEGINNING WITH U20601957 (DRY SYSTEM)
TROUBLE

PROBABLE CAUSE

REMEDY

HIGH SUCTION GAGE READINGS. Gyros function normally-relief
valve screen clogged, relief
valve malfunction.

Check screen, then valve. Compare gage readings with new gage.
Clean screen, reset valve. Replace gage.

NORMAL SUCTION GAGE
READING, SLUGGISH OR

Instrument air filters clogged.

Clean or replace filter as
necessary.

Leaks or restriction between
instruments and relief valve,
relief valve out of adjustment,
defective pump.

Check lines for leaks, disconnect
and test pump. Repair or replace
lines, adjust or replace relief
valve, repair or replace pump.

Central air filter dirty.

Clean or replace filter as
necessary

ERRATIC GYRO RESPONSE.

LOW SUCTION GAGE
READINGS.

16-12B

Change 3

16-25A.

TROUBLE SHOOTING--BEGINNING WITH U20601957 DRY SYSTEM (Cont)
TROUBLE

SUCTION GAGE FLUCTUATES.

16-26.

PROBABLE CAUSE
Defective gage or sticking relief
valve.

REMEDY
Check suction with test gage.
Replace gage. Clean sticking
valve with Stoddard solvent.
Blow dry and test. If valve
sticks after cleaning, replace
valve.

TROUBLE SHOOTING--GYROS.

TROUBLE
HORIZON BAR FAILS TO
RESPOND.

HORIZON BAR DOES NOT
SETTLE.

HORIZON BAR OSCILLATES
OR VIBRATES EXCESSIVELY.

PROBABLE CAUSE

REMEDY

Central filter dirty.

Check filter.
filter.

Suction relief valve improperly
adjusted.

Adjust or replace relief valve.

Faulty suction gage.

Substitute known-good suction
gage and check gyro response.
Replace suction gage.

Vacuum pump failure.

Check pump.

Vacuum line kinked or leaking.

Check lines for damage and leaks.
Repair or replace damaged lines.
tighten connections.

Defective mechanism.

Substitute known-good gyro and
check indication. Replace instrument.

Insufficient vacuum.

Adjust or replace relief valve.

Excessive vibration.

Check panel shock-mounts.
Replace defective shock-mounts.

Central filter dirty.

Check filter.
filter.

Suction relief valve improperly adjusted.

Adjust or replace relief valve.

Faulty suction gage.

Substitute known-good suction
gage and check gyro indication.
Replace suction gage.

Defective mechanism.

Substitute known-good gyro and
check indication. Replace instrument.

Excessive vibration.

Check panel shock-mounts. Replace defective shock-mounts.

Clean or replace

Replace pump.

Clean or replace

Change 1

16-13

16-26.

TROUBLE SHOOTING--GYROS.
TROUBLE

EXCESSIVE DRIFT IN
EITHER DIRECTION.

(Cont).
PROBABLE CAUSE

REMEDY

Clean or replace

Central air filter dirty.

Check filter.
filter.

Low vacuum, relief valve
improperly adjusted.

Adjust or replace relief valve.

Faulty suction gage.

Substitute known-good suction
gage and check gyro indication.
Replace suction gage.

Vacuum pump failure.

Check pump.

Vacuum line kinked or
leaking.

Check lines for damage and
leaks. Repair or replace damaged lines, tighten connections.

Replace pump.

DIAL SPINS IN ONE
DIRECTION CONTINU-

Operating limits have been
exceeded.

Replace instrument.

OUSLY.

Defective mechanism.

Substitute known-good gyro
and check indication. Replace
instrument.

16-27.

TROUBLE SHOOTING--VACUUM PUMP (Wet System)

TROUBLE

PROBABLE CAUSE

REMEDY

Damaged engine drive seal.

Replace gasket.

Oil separator clogged, oil
return line obstructed, excessive oil flow through pump.

Clean oil separator with Stoddard
solvent, then blow dry. Blow out
lines. If pump oil consumption is
excessive, replace oil metering
pin in pump.

HIGH SUCTION.

Suction relief valve
screen clogged.

Clean or replace screen.

LOW SUCTION.

Relief valve leaking.

Replace relief valve.

Vacuum pump failure.

Replace vacuum pump.

EXCESSIVE OIL IN DISCHARGE.

16-27A.

TROUBLE SHOOTING-- VACUUM PUMP (Dry System)

TROUBLE
OIL IN DISCHARGE.

16-14

Change 1

PROBABLE CAUSE
Damaged pump drive seal.

REMEDY
Replace gasket.

© 16-27A.

TROUBLE SHOOTING-VACUUM PUMP (Wet System)
TROUBLE

HIGH SUCTION.

LOW SUCTION.

PROBABLE CAUSE
Suction relief valve
screen clogged.
elief valve leaking.

16-28. REMOVAL AND INSTALLATION OF COMPONENTS. Through Aircraft Serial U20601956 the
varicus components of the vacuum system are secured by conventional clamps, mounting screws and
nuts. To remove a component, remove mounting
screws and disconnect inlet anrd discharge lines.
When replacing a vacuum system component. ensure
connections are made correctly. Use thread lubricant sparingly and only on male threads. Avoid overtightening connections. Before reinstalling a vacuum
pump, probe oil passages in pump and engine, to
make sure they are open. Place mounting pad gasket
in position over studs and ensure it does not block oil
passages. Coat pump drive splines lightly with a
high-temperature grease such as Dow Silicone #30
Dow-Corning Co., Midland, Mich.). ™fterinstalling pump, before connecting plumbing, start engine
and hold a piece of paper over pump discharge to
check for proper lubrication. Proper oil flow through
pump is one to ifour fluid ouncrs per hour.
16-28A. RIEMOVALI
ANID INSfTALLATION OF COMPONENTS. Beginning with U20 6 01957 the various
comiponents of the vacuum system are secured by conveonionat clamps, rmounting screws and nuts. To reimve a component, remove mounting screws and discomnnec in'lv aji :tischariLe lines. Cap open lines and
liltiig to pieveni. dir: fr.ri lnlerilng the system.
U'l-en repiacing a vacuum systemi component, ensure
connecimins are miade correctly. Use no lubricanls
on anv components when assemhlng a dry vacuum
sysiemn. Avoid over-tightening connections. Before
installing the vacuum pump, place mounting pad gasket in position over studs. Be sure all lines and fittings are open and caps are removed.

REMEDY
Clean or replace screen.

Replace relief valve.

Vacuum pump failure.

SHOP NOTES:

(Cont)

Replace vacuum pump.

16-29. CLEANING. Low pressure, dry compressed
air should be used in cleaning vacuum system components. The suction relief valve should be washed
with Stoddard solvent then dried with low-pressure
air. Refer to Section 2 for central air filter. Check
hose for collapsed inner liners as well as external
damage.
CAUTION
Never apply compressed air to lines or components installed in aircraft. The excessive
pressures will damage gyros. If an obstructed line is to be blown out, disconnect at both
ends and blow from instrument panel out.
16-30. VACUUM RELIEF VALVE ADJUSTMENT.
A suction gage reading of 5. 3 inches of mercury is
desirable for gyro instruments. However, a range
of 4.6 to 5.4 inches of mercury is acceptable. To
adjust relief valve, remove control air filter. run
engine to 2200 rpm on ground and adjust relief valve
to 5.3 ± .1 inches of mercury.
CAUTIONI
Do not exceed maximum engine temperature.
N
The relief valve on turbocharged aircraft
is alitude compensated by an internal aneroid. Operation of the compensating
mechanism is automatic. Standard relief
valve adjustment applies to the compensated relief valve.
Be sure filter element is clean before installing.
reading drops noticeably,

If

install new filter element.

Change 1

16-14A/16-14B(blank)

NOTE

16-30. VACUUM RELIEF VALVE ADJUSTMENT.
A suction gage reading of 5.3 inches of mercury is
desirable for gyro instruments. However, a range
of 4.6 to 5.4 inches of mercury is acceptable. To
adjust relief valve, remove control air filter, run
engine to 2200 rpm on ground and adjust relief valve
to 5.3 ± . 1 inches of mercury.

Before replacing a tachometer cable in hous
Ing, coat lower two thirds with AC Type ST640 speedometer cable grease or Lubriplate
No. 110. Insert cable in housing as far as
possible, then slowly rotate to make sure it
is seated in engine fitting. Insert cable in
tachometer, making sure it is seated in drive
shaft, then reconnect housing and torque to
50 pound-inches (at instrument).

CAUTION
Do not exceed maximum engine temperature.
Be sure filter element is clean before installing. If
reading drops noticeably, install new filter element.
16-31.

ENGINE INDICATORS.

16-32.

TACHOMETER.

16-33. DESCRIPTION. The tachometer is a mechanical indicator driven at half crankshaft speed by a
flexible shaft. Most tachometer difficulities will be
found in the drive-shaft. To function properly, the
shaft housing must be free of kinks, dents and sharp
bends. There should be no bend on a radius shorter
than six inches and no bend within three inches of
either terminal. If a tachometer is noisy or pointer
oscillates. check cable housing for kinks, sharp
bends and damage. Disconnect cable at tachometer
and pull it out of housing. Check cable for worn
spots. breaks and kinks.
16-36.

16-34. MANIFOLD PRESSURE/FUEL FLOW INDICATOR.
15-35. DESCRIPTION. The manifold pressure and
fuel flow indicators are in one instrument case.
However, each instrument operates independently.
The manifold pressure gage is a barometric instrument which indicates absolute pressure in the intake
manifold inches of mercury. The fuel flow indicator
is a pressure instrument calibrated in gallons per
hour, indicating approximate gallons of fuel metered
per hour to the engine. Pressure for operating the
indicator is obtained through a hose from the fuel
manifold valve. The fuel flow indicator is vented to
atmospheric pressure with standard engines and to
turbocharger outlet pressure on turbocharged engines.

TROUBLE SHOOTING -- FUEL FLOW INDICATOR.
TROUBLE

DOES NOT REGISTER.

POINTER FAILS TO RETURN
TO ZERO.

INCORRECT OR ERRATIC
READING.

PROBABLE CAUSE

REMEDY

Pressure line clogged.

Blow out line.

Pressure line broken.

Repair or replace damaged line.

Fractured bellows or
damaged mechanism.

Replace instrument.

Clogged snubber orifice.

Replace instrument.

Pointer loose on staff.

Replace instrument.

Foreign matter in line.

Blow out line.

Clogged snubber orifice.

Replace instrument.

Damaged bellows or
mechanism.

Replace instrument.

Damaged or dirty mechanism.

Replace instrument.

Pointer bent, rubbing on dial
or glass.

Replace instrument.

Leak or partial obstruction
in pressure or vent line.

Blow out dirty line, repair
or tighten loose connections.

16-15

16-37.

TROUBLE SHOOTING -- MANIFOLD PRESSURE INDICATOR.
TROUBLE

PROBABLE CAUSE

EXCESSIVE ERROR AT EXISTING BAROMETRIC PRESSURE.

REMEDY

Pointer shifted.

Replace instrument.

Leak in vacuum bellows.

Replace instrument.

Loose pointer.

Replace instrument.

Leak in pressure line.

Repair or replace damaged
line, tighten connections.

Condensate or fuel in line.

Blow out line.

Excessive internal friction.

Replace instrument.

Rocker shaft screws tight.

Replace instrument.

Link springs too tight.

Replace Instrument.

Dirty pivot bearings.

Replace instrument.

Defective mechanism.

Replace instrument.

Leak in pressure line.

Repair or replace damaged
line, tighten connections.

Foreign matter in line.

Blow out line.

Damping needle dirty.

Replace Instrument.

Leak in pressure line.

Repair or replace damaged line,
tighten connections.

Tight rocker pivot bearings.

Replace instrument.

Excessive vibration.

Tighten mounting screws.

IMPROPER CALIBRATION.

Faulty mechanism.

Replace Instrument.

NO POINTER MOVEMENT.

Faulty mechanism.

Replace instrument.

Broken pressure line.

Repair or replace damaged
line.

JERKY MOVEMENT OF
POINTER.

SLUGGISH OPERATION OF
POINTER.

EXCESSIVE POINTER VIBRATION.

16-16

_

16-38. CYLINDER HEAD TEMPERATURE GAGE.
16-39. DESCRIPTION. The temperature sending unit
regulates power through the cylinder head temperature
gage. The gage and sending unit require little or no
maintenance other than cleaning, making sure the lead is
properly supported, and all connections are clean and
properly insulated. The Rochester and Stewart Warner
gages are connected the same, but the Rochester gage does
16-40.

not have a calibration pot and cannot be adjusted. Refer to
Table 2 on page 16-18C/D when trouble shooting the
cylinder head temperature gage.
NOTE
A Cylinder Head Temperature Gage Calibration Unit, (SK182-43) is available and may be
ordered through the Cessna Supply Division.

TROUBLE SHOOTING.
TROUBLE

GAGE INOPERATIVE.

PROBABLE CAUSE

REMEDY

No current to circuit.

Repair electrical circuit.

Defective gage, bulb or
circuit.

Repair or replace defective
items.

GAGE FLUCTUATES
RAPIDLY.

Loose or broken wire permitting alternate make and
break of gage circuit.

Repair or replace defective
wire.

GAGE READS TOO HIGH
ON SCALE.

High voltage.

Check "A" terminal.

Gage off calibration.

Recalibrate or replace gage.

Low voltage.

Check voltage supply and
"D" terminal.

Gage off calibration.

Recalibrate or replace gage.

Break in bulb.

Replace bulb.

Break in bulb lead.

Replace bulb.

Internal break in gage.

Replace gage.

Defective gage mechanism.

Replace gage.

Incorrect calibration.

Recalibrate.

GAGE READS TOO LOW
ON SCALE.

GAGE READS OFF SCALE
AT HIGH END.

OBVIOUSLY INCORRECT
READING.

16-41.

OIL PRESSURE GAGE.

16-42. DESCRIPTION. The Bourdon tube-type oil
pressure gage is a direct-reading instrument, operated by a pressure pickup line connected to the engine

Temporary Revision 3 - Oct 3/94

main oil gallery. The oil pressure line from the instrument to the engine should be filled with kerosene,
especially during cold weather operation, to attain
an immediate oil indication.

16-17

16-43.

TROUBLE SHOOTING.
TROUBLE

GAGE DOES NOT REGISTER.

PROBABLE CAUSE

REMEDY

Pressure line clogged.

Check line for obstructions. Clean
line.

Pressure line broken.

Check line for leaks and damage.
Repair or replace damaged line.

Fractured Bourdon tube.

Replace instrument.

Gage pointer loose on staff.

Replace instrument.

Damaged gage movement.

Replace instrument.

Foreign matter in line.

Check line for obstructions.
Clean line.

Foreign matter in Bourdon
tube.

Replace instrument.

Bourdon tube stretched.

Replace instrument.

GAGE DOES NOT REGISTER
PROPERLY.

Faulty mechanism.

Replace instrument.

GAGE HAS ERRATIC
OPERATION.

Worn or bent movement.

Replace instrument.

Foreign matter in Bourdon
tube.

Replace instrument.

Dirty or corroded movement.

Replace instrument.

Pointer bent and rubbing on
dial, dial screw or glass.

Replace instrument.

Leak in pressure line.

Check line for leaks and damage. Repair or replace
damaged line.

GAGE POINTER FAILS
TO RETURN TO ZERO.

16-44. OIL TEMPERATURE GAGE.
16-45. DESCRIPTION. On some airplanes, the oil temperature gage is a Bourdon tubetype pressure instrument
connected by armored capillary tubing to a temperature
bulb in the engine The temperature bulb, capillary tube
and gage are filled with fluid and sealed. Expansion and
contraction of fluid in the bulb with temperature changes
operates the gage. Checking capillary tube for damage and
fittings for security is the only maintenance required. Since
the tubes inside diameter is small, small dents and kinks,
which would be acceptable in larger tubing, may partially
or completely close off the capillary, making the gage
inoperative. Some airplanes are equipped with gages that
are electrically actuated and are not adjustable.Table 1 on
page 16-18B when trouble shooting the cylinder head
temperature gage.
16-46.

FUEL QUANTITY INDICATING SYSTEM.

operated variable-resistance transmitter in each
fuel tank. The full position of float produces a minimum resistance through transmitter, permitting
maximum current flow through the fuel quantity indicator and maximum pointer deflection. As fuel level
is lowered, resistance in transmitter is increased,
producing a decreased current flow through fuel quantity indicator and a smaller pointer deflection. Beginning with Serial U206-01573, a heat sink assembly
(Voltage Regulator) is incorporated into the fuel quantity indicating system of aircraft equipped with a 24volt system. The unit is mounted on top of the glove
box thru U20602199 and is located under the glove
box beginning with U20602200. The unit converts 28volt current flow from the bus to a 14-volt current
flow to the fuel quantity indicators and transmitters.
Refer to the 24-volt part of Section 20 in this Service
Manual for a schematic wiring diagram of the Heat
Sink Assembly.

16-47. DESCRIPTION. The magnetic type fuel quantity indicators are used in conjunction with a float16-18

Change 1

Temporary Revision 3 - Oct 3/94

16-48.

TROUBLE SHOOTING.
TROUBLE

PROBABLE CAUSE

REMEDY

FAILURE TO INDICATE.

No power to indicator or transmitter. (Pointer stays below E. )

Check fuse and inspect for open
circuit. Replace fuse, repair
or replace defective wire.

Grounded wire.
above F.)

Check for partial ground between
transmitter and gage. Repair or
replace defective wire.

OFF CALIBRATION.

STICKY OR SLUGGISH
INDICATOR OPERATION.

ERRATIC READINGS.

16-49.

(Pointer stays

Low voltage.

Check voltage at indicator.
Correct voltage.

Defective indicator.

Substitute known-good indicator.
Replace indicator.

Defective indicator.

Substitute known-good indicator.
Replace indicator.

Defective transmitter.

Substitute known-good transmitter.
Recalibrate or replace.

Low or high voltage.

Check voltage at indicator.
Correct voltage.

Defective indicator.

Substitute known-good indicator.
Replace indicator.

Low voltage.

Check voltage at indicator.
Correct voltag

Loose or broken wiring on
indicator or transmitter.

Inspect circuit wiring.
Repair or replace defective wire.

Defective indicator or transmitter.

Substitute known-good component.
Replace indicator or transmitter.

Defective master switch.

Replace switch.

TRANSMITTER ADJUSTMENT.
(Refer to page 16-18B).

16-49C. REMOVAL AND INSTALLATION FUEL
QUANTITY TRANSMITTERS. (Refer to Section 13,
figure 13-5.) Observe precautions of Section 13-3
when working-with fuel components.
a. Drain fuel from cell.
b. Remove wing root fairing.
c. Disconnect electrical lead and ground strap from
transmitter.
d. Remove screws through transmitter and wing
root rib, and remove transmitter.
Temporary Revision 3 - Oct 3/94

Change 1

16-18A

TRANSMITTER ADJUSTMENT.

16-49.

WARNING

Using the following fuel transmitter calibration procedure on components other than the originally
installed (Stewart Warner) components will result in a faulty fuel quantity reading.
16-49A. STEWART WARNER GAGE TRANSMITTER CALIBRATION. Chances of transmitter calibration
changing in normal service is remote; however, it is possible that float arm or float arm stops may become
bent if transmitter is removed from cell. Transmitter calibration is obtained by adjusting float travel. Float
travel is limited by float arm stops.
WARNING

Use extreme caution while working with electrical components of the fuel system. The possibility of
electrical sparks around an "empty" fuel cell creates a hazardous situation.
Before installing transmitter, attach electrical wires and place master switch in "ON" position. Allow float
arm to rest against lower float arm stop and read indicator. The pointer should be on E (empty) position.
Adjust the float arm against lower stop so pointer indicator is on E. Raise float until arm is against upper
stop and adjust upper stop to permit indicator pointer to be on F (full). Install transmitter in accordance with
paragraph 16-49C.
16-49B. ROCHESTER GAGE TRANSMITTER. Do not attempt to adjust float arm or stop. No adjustment is
allowed.
Table 1
NOTE
Select the oil temperature sending unit part number that is used in your aircraft

from the left column and the temperature from the column headings. Read the ohms
value under the appropriate temperature column.
72°F

120°F

165°F

220°F

250°F

Part Number

Type

S1630-1

Oil Temp

S1630-3

Oil Temp

620.0

52.4

S1630-4

Oil Temp

620.0

52.4

S1630-5

Oil Temp

S2335-1

Oil Temp

16-18B

46.4

192.0
990.0

34.0

Temporary Revision 3 - Oct 3/94

16-49C.

CYLINDER HEAD TEMPERATURE INDICATING SYSTEM RESISTANCE TABLE 2
The following table is provided to assist in the troubleshooting the cylinder head temperature indicating
system components.
Select the cylinder head temperature sending unit part number that is used in your airplane from the
left column and the temperature from the column headings. Read the ohms value under the
appropriate temperature column.
Part Number
S1372-1
S1372-2
S1372-3
S1372-4
S2334-3
S2334-4

16-49D.

200°F

Type
CHT
CHT
CHT
CHT
CHT
CHT

220°F
310.0
310.0

450°F
34.8
34.8
113.0
113.0

745.0
745.0

475°F

38.0
38.0

FUEL QUANTITY INDICATING SYSTEM OPERATIONAL TEST
WARNING:

REMOVE ALL IGNITION SOURCES FROM THE AIRPLANE AND VAPOR HAZARD
AREA. SOME TYPICAL EXAMPLES OF IGNITION SOURCES ARE STATIC
ELECTRICITY, ELECTRICAL POWERED EQUIPMENT (TOOLS OR ELECTRONIC
TEST EQUIPMENT - BOTH INSTALLED ON THE AIRPLANE AND GROUND
SUPPORT EQUIPMENT), SMOKING AND SPARKS FROM METAL TOOLS.

WARNING:

OBSERVE ALL STANDARD FUEL SYSTEM FIRE AND SAFETY PRACTICES.

1. Disconnect all electrical power from the airplane. Attach maintenance warning tags to the battery
connector and external power receptacle stating:
DO NOT CONNECT ELECTRICAL POWER, MAINTENANCE IN PROGRESS
2.

Electrically ground the airplane.

3.

Level the airplane and drain all fuel from wing fuel tanks.

4.

Gain access to each fuel transmitter float arm and actuate the arm through the transmitter's full
range of travel.
A.

Ensure the transmitter float arm moves freely and consistently through this range of travel.
Replace any transmitter that does not move freely or consistently.

WARNING: USE EXTREME CAUTION WHILE WORKING WITH ELECTRICAL COMPONENTS
OF THE FUEL SYSTEM. THE POSSIBILITY OF ELECTRICAL SPARKS AROUND
AN "EMPTY" FUEL CELL CREATES A HAZARDOUS SITUATION.
B. While the transmitter float arm is being actuated, apply airplane battery electrical power as
required to ensure that the fuel quantity indicator follows the movement of the transmitter float
arm. If this does not occur, troubleshoot, repair and/or replace components as required until
the results are achieved as stated.
NOTE: Stewart Warner fuel quantity indicating systems can be adjusted. Refer to paragraph
16-49A for instructions for adjusting Stewart Warner fuel indicating systems.
Rochester fuel quantity indicating system components are not adjustable, only
component replacement or standard electrical wiring system maintenance practices
are

Temporary Revision Number 5
6 January 2003

permitted.

© 2003 CESSNA AIRCRAFT COMPANY

16-18C

5.

With the fuel selector valve in the "OFF" position, add unusable fuel to each fuel tank.

6.

Apply electrical power as required to verify the fuel quantity indicator indicates "EMPTY".
A.

If "EMPTY" is not indicated, adjust, troubleshoot, repair and/or replace fuel indicating components
as required until the "EMPTY" indication is achieved.
NOTE: Stewart Warner fuel quantity indicating systems can be adjusted. Refer to paragraph 1649A for instructions for adjusting Stewart Warner fuel indicating systems.
Rochester fuel quantity indicating system components are not adjustable, only component
replacement or standard electrical wiring system maintenance practices are permitted.

7.

Fill tanks to capacity, apply electrical power as required and verify fuel quantity indicator indicates
"FULL".
A. If "FULL" is not indicated, adjust, troubleshoot, repair and/or replace fuel indicating components as
required until the "FULL" indication is achieved.
NOTE: Stewart Warner fuel quantity indicating systems can be adjusted. Refer to paragraph 1649A for instructions for adjusting Stewart Warner fuel indicating systems.
Rochester fuel quantity indicating system components are not adjustable, only component
replacement or standard electrical wiring system maintenance practices are permitted.

8.

16-1118D

Install any items and/or equipment removed to accomplish this procedure, remove maintenance
warning tags and connect the airplane battery.

2003 CESSNA AIRCRAFT COMPANY

Temporary Revision Number 5
6 January 2003

.

HEAT SINK ASSEMBLY
GLOVE BOX COVER

Figure 16-6. Heat Sink Assembly (Voltage Regulator) Installation
e. Install transmitter by reversing preceding steps.
No gasket paste should be used.
f. Fill fuel cell. Check for leaks and correct fuel
quantity indication.
NOTE
Be sure grounding is secure and in accordance
with figure 13-5.

16-49B- REMOVAL AND INSTALLATION HEAT
SINK. (Refer to figure 16-6.)
a. Turn off master switch or disconnect battery
leads.

b. Disconnect 3 wires from heat sink assembly and
tag for identification.
c. Remove nuts, screws and washers attaching unit
to glove box and remove the unit.
d. Reverse preceding steps to install the heat sink
unit.
16-50.

HOURMETER.

16-51. DESCRIPTION. The hourmeter is electrically operated instrument, actuated by a pressure
switch in the oil pressure gage line. Electrical
power is supplied through a one-amp fuse from the
electrical clock circuit, and therefore will operate
independent of master switch.

SHOP NOTES:

Change 1

16-19

16-52.

ECONOMY MIXTURE INDICATOR.

16-53. DESCRIPTION. The economy mixture indicator is an exhaust gas temperature (EGT) sensing
device which is used to aid pilot in selecting most
16-54.

desirable fuel-air mixture for cruising flight at
less than 75% power. Exhaust gas temperature (EGT)
varies with ratio of fuel-to-air mixture entering
engine cylinders. Refer to Owner's Manual for
operating procedure of system.

TROUBLE SHOOTING.
PROBABLE CAUSE

TROUBLE

REMEDY

GAGE INOPERATIVE.

Defective gage, probe or
circuit.

Repair or replace defective
part.

INCORRECT READING.

Indicator needs calibrating.

Calibrate indicator in accordance
with paragraph 15-56.

FLUCTUATING READING.

Loose, frayed or broken
lead, permitting alternate
make and break of circuit.

Tighten connections and repair or replace defective
leads.

16-55. CALIBRATION. A potentiometer adjustment
screw is provided behind the plastic cap at the back
of the instrument for calibration. This adjustment
screw is used to position the pointer over the reference increment line (4/5 of scale) at peak EGT. Establish 65% power in level flight, then carefully lean
the mixture to peak EGT. After the pointer has peaked,
using the adjustment screw, position pointer over the
reference increment line (4/5 of scale).
NOTE
This setting will provide relative temperature indications for normal cruise power
settings within range of the instrument.
Turning the screw clockwise increases the meter
reading and counterclockwise decreases the meter
reading. There is a stop in each direction and damage can occur if too much torque is applied against
stops. Approximately 600°F total adjustment is provided. The adjustable yellow pointer on the face of
the instrument is a reference pointer only.
16-56. REMOVAL AND INSTALLATION. Removal
of the indicator is accomplished by removing the
mounting screws and disconnecting the leads. Tag
leads to facilitate installation. The thermocouple
probe is secured to the exhaust stack with a clamp.
When installing probe, tighten clamp to 45 poundinches and safety as required.

16-20

Change 1

16-57.

MAGNETIC COMPASS.

16-58. DESCRIPTION. The magnetic compass is
liquid-filled, with expansion provisions to compensate for temperature changes. It is equipped with
compensating magnets adjustable from the front of
the case. The compass is internally lighted, controlled by the panel lights rheostat. No maintenance
is required on the compass except an occasional
check on a compass rose and replacement of the lamp.
The compass mount is attached by three screws to a
base plate which is bonded to the windshield with
methylene chloride. A tube containing the compass
light wires is attached to the metal strip at the top of
the windshield. Removal of the compass is accomplished by removing the screw at the forward end of
the compass mount, unfastening the metal strip at the
top of the windshield and cutting the two wire splices.
Removal of the compass mount is accomplished by
removing the outside air temperature probe and removing the three screws attaching mount to the base
plate. Access to the inner screw is gained through a
hole in the bottom of mount, through which a thin
screwdriver may be inserted. When installing the
compass, it will be necessary to splice the compass
light wires.
16-59.

STALL WARNING HORN AND TRANSMITTER.

16-60. DESCRIPTION. The stall warning horn is
mounted on the glove box. It is electrically operated

and controlled by a stall warning transmitter mounted on leading edge of left wing. For further information on warning horn and transmitter, refer to
Section 17.
16-61.

TURN-AND-SLIP INDICATOR.

16-63.

TROUBLE SHOOTING.

TROUBLE
INDICATOR POINTER FAILS TO
RESPOND.

16-62. DESCRIPTION. The turn-and-slip indicator
is operated by the aircraft electrical system and
operates ONLY when the master switch is on. Its
circuit is protected by an automatically-resetting
circuit breaker.

PROBABLE CAUSE

REMEDY

Automatic resetting circuit
breaker defective.

Check circuit breaker.
Replace circuit breaker.

Master switch "OFF" or
switch defective.

Check switch "ON."
defective switch.

Broken or grounded lead to
indicator.

Check circuit wiring. Repair
or replace defective wiring.

Indicator not grounded.

Check ground wire. Repair
or replace defective wire.

Defective mechanism.

Replace instrument.

Defective mechanism.

Replace instrument.

Low voltage.

Check voltage at indicator.
Correct voltage.

POINTER DOES NOT INDICATE
PROPER TURN.

Defective mechanism.

Replace instrument.

HAND DOES NOT SIT
ON ZERO

Gimbal and rotorout of balance.

Replace instrument

Hand incorrectly sits on rod.

Replace instrument

Sensitivity spring adjustment
pulls hand off zero.

Replace instrument.

Oil in indicator becomes too
thick.

Replace instrument

Insufficient bearing end play.

Replace instrument.

Low voltage.

Check voltage at indicator.
Correct voltage.

High voltage.

Check voltage at indicator.
Correct voltage.

Loose or defective rotor
bearings.

Replace instrument.

HAND SLUGGISH IN
RETURNING TO ZERO.

IN COLD TEMPERATURES,
HAND FAILS TO RESPOND
OR IS SLUGGISH.

NOISY GYRO.

Replace

16-21

16-64.

TURN COORDINATOR.

16-65. DESCRIPTION. The turn coordinator is an
electrically operated, gyroscopic, roll-rate turn
indicator. Its gyro simultaneously senses rate of
16-66.

motion roll and yaw axes which is projected on a
single indicator. The gyro is a non-tumbling type requiring no caging mechanism and incorporates an
a. c. brushless spin motor with a solid state inverter.

TROUBLE SHOOTING.
TROUBLE

PROBABLE CAUSE

INDICATOR DOES NOT
RETURN TO CENTER.

Friction caused by contamination
in the indicator damping.

Replace instrument.

Friction in gimbal assembly.

Replace instrument.

Low voltage.

Measure voltage at instrument.
Correct voltage.

Inverter frequency changed.

Replace instrument.

NOISY MOTOR.

Faulty bearings.

Replace instrument.

ROTOR DOES NOT START.

Faulty electrical connection.

Check continuity and voltage.
Correct voltage or replace
faulty wire.

Inverter malfunctioning.

Replace instrument.

Motor shorted.

Replace instrument.

Bearings frozen.

Replace instrument.

Oil in indicator becomes
too thick.

Replace instrument.

Insufficient bearing end play.

Replace instrument.

Low voltage.

Check voltage at instrument.
Correct voltage.

High voltage.

Check voltage to instrument.
Correct voltage.

Loose or defective rotor
bearings.

Replace instrument.

DOES NOT INDICATE A
STANDARD RATE TURN
(TOO SLOW).

IN COLD TEMPERATURES,
HAND FAILS TO RESPOND
OR IS SLUGGISH.

NOISY GYRO.

16-22

REMEDY

1. Windshield

c\t
\

M//

y

4 f \\

^/

^
'/~

/



,:,

^^

39

/

* BEGINNING WITH AIRCRAFT
SERIAL U20601875

17

/

BEGINNING WITH 1971 MODELS

1

V
,4

w.s. 118. 00

f

9
1

.19

14

\

"21\

r

BEGINNING WITH

~~19^^~I

1^

'L

1973 MODELS

7
THRU 1972 MODELS

1.
2.
3.
4.
5.
6.
7.
8.

Electrical Leads
Cap
Washer
Insulated Washer
Spring
Insulator
Housing - Plug
Housing -Cap
Figure 17-8.

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

Wing Tip
Wing Navigation Light
Spacer
Flash Tube Assembly
Lens
Screw
Lens Retainer
Bulb
Seal

18.
19.
20.
21.
22.
23.
24.
25.

Bracket
Nutplate
Bolt
Power Supply
Inspection Plate
Rear Spar
Wing Tip Rib
Gasket

Navigation and Anti-Collision Strobe Lights Installation (Sheet 2 of 2)
Change 1

17-31

* THRU 1972 MODELS
* BEGINNING WITH 1973 MODELS

A

6

4. ^--i4\

^s

//

IrS

*

n410

a

e^

-

,

>

,^^

'

-at,.>f

DetailB

3

~~~~~~~~~~~~~~~~~14

^

^

ICAUTIONI

//^*^^'

When inserting lamp into socket
always use a handkerchief or a
tissue to prevent getting fingerprints on the lamp.

_I~~~. ^~

NOTE

Detail A
Fingerprints on lamp may shorten the life of the lamp.
1.
2.
3.
4.
5.
6.

Dome
Gasket
Lamp
Screw
Baffle
Clamp Assembly

7.
8.
9.
10.
11.
12.
Figure 17-9.

17-32

Change 1

4

//~18

V

12^.

-. '2C

1
lF
13i
:7=*

\04

,

-'/

Socket Assembly
Nutplate
Tip Assembly - Fin
Spacer
Flasher Assembly
Fin Assembly

_

13.
14.
15.
16.
17.
18.

Flashing Beacon Light Installation

Housing - Cap
Housing - Plug
Plate
Stabilizer Skin - Upper
Resistor
Washer

B

A

AIRCRAFT SERIALS THRU P20600648
AND U206-1235 THRU P20601587

10
2

4

q

6

><6V

-6

\ -1'

DETAIL A
TYPICAL
INSTALLATION

1.
2.
3.

22

10

Light Fitting Assembly
Nut
Light Assembly

Figure 17-10.

88

DETAIL B

4.
5.
6.
7.

8

DETAIL C

Retainer
Washer
Bracket
Gasket

8.
9.
10.

Cover
Screw
Bulb

Instrument Panel Glare Shield Light Installation (Sheet 1 of 2)
17-33

AIRCRAFT SERIAL U20601588
THRU U20601700

AIRCRAFT SERIAL U20601701
THRU U20601874

01 WITH 5BEGINNING
AIRCRAFT
~
SERIAL U20601875

I,

3

2
2

396

2

-

X

4

"-- ---S
12 VOLT

24 VOLT

Detail A

Detail A

1. Reflector
2. Lamp

5. Screw
6. Nut

3. Lamp Socket
4. Housing

7.
8.

Figure 17-10.
17-34

Tinnerman Screw
Tinnerman Nut
Instrument Panel Glare Shield Light Installation (Sheet 2 of 2)

7

1

5

1.

Screw

2. Washer
3.
4.

Transistor
Mica Washer

5.

Housing - Socket

6.
7.

Heat Sink
Mounting Bracket

Detail A
Figure 17-11.

Transistorized Light Dimming Installation

and the comfort control panel. The ac voltage required to drive the "EL" panels is supplied by a small
inverta-pak (power supply) located behind the instrument panel. The intensity of the 'EL" panel lighting
is controlled by a rheostat located on the instrument
panel. Beginning with aircraft serials P20600635
and U20601493 a resistor is installed ahead of the
dimming EL rheostat as a lood for the AC output of
the E inverter. Due to heat dissipation, the resistor
must be kept away from the wire bundle. Refer to
figure 17-1 and 17-13.
17-81.

PEDESTAL LIGHTS .

17-82. DESCRIPTION. The pedestal lights consist
of two post type lights mounted on the pedestal to
illuminate the rudder and elevator trim controls.

The pedestal lights are controlled by the instrument
light rheostat.
17-83. REMOVAL AND INSTALLATION. For removal and replacement of the pedestal lamp, slide
the cap and lens assembly from the base. Slide the
lamp from the socket and replace.
17-84.

INSTRUMENT POST LIGHTING.

17-85. DESCRIPTION. Individual post lighting may
be installed as optional equipment to provide for nonglare instrument lighting. The post light consists of
a cap and a clear lamp assembly with a tinted lens.
The intensity of the instrument post lights is controlled by the radio light dimming rheostat located
on the switch panel.

Change 1

17-35

NOTE
Adjust the overhead map light so that the forward edge of the lighted area is 3.0 (±1.0)
inches aft of the control wheel (when full forward).

1.
2.
3.
4.
5.
6.

Panel Light Bracket
Nutplate
Washer
Spacer
Spring
Tinnerman Nut

7.
8.
9.
10.
11.

Figure 17-12.
17 -36
17-36

Panel Light Housing
Lamp
Clip
Slide Cover
Adjustment Screw

Overhead Console Installation

12.
13.
14.
15.
16.
17.

Screw
Slide Knob
Panel Light Cover
Lamp Socket
Grommet
Nut

A
A-

2

BEGINNING WITH U2060175

1. Nut
2. Inverta-pak, Power Supply
3. Washer
4. Screw

Figure 17-13.

Detail B
THRU U20601874

W

Electroluminescent Panel Inverta-pak Power Supply
Change 2

17-37

17-86. REMOVAL AND INSTALLATION. For removal and replacement of the instrument post lamps,
slide the cap and the lens assembly from the base.
Slide the lamp from the socket and replace.
17-87.

c. Detach wires from the terminal strip along the
edge of the circuit board. Note the connection for
reference when replacing the board.
d. To install the control wheel map light, reverse
the procedure.

COURTESY LIGHTS.
NOTE

17-88. DESCRIPTION. The lights consist of one
light located on the underside of each wing to provide ground lighting around the cabin area. The
courtesy lights have clear lens and are controlled
by a single slide switch labeled, "Utility Lights,"
located on the left rear door post. The switch also
operates the dome lights thru 1972 Models.
17-89. REMOVAL AND INSTALLATION. Refer to
Figure 17-14 for removal and installation.
17-90. INTERIOR LIGHTING. Thru 1972 Models
the cabin interior is illuminated by two dome lights,
one dome light on each side of the aft cabin. The
dome lights are controlled by a single slide switch
labeled "Utility Lights," located on the left door post.
The switch also operates the courtesy lights. Beginning with 1973 Models a single dome light is installed overhead center aft of the rear spar. The
light is controlled by a rocker switch on the assembly.
17-91. REMOVAL AND INSTALLATION. Thru 1972
models for removal and replacement of dome lamps,
pry light assembly out of retainer then pry socket
out of light assembly. Twist the bayonet type lamp
from the socket and replace. Beginning with 1973
models the lens snap out for access to the lamp.
17-92.

17-95. REMOVAL AND INSTALLATION (AIRCRAFT
U20601445 THRU U20601700) (Refer to Figure 17-15.)
a. Rotate the control wheel 90° to the left to gain
access to the underside of the control wheel.
b. Remove two screws and nuts holding map light
assembly to control wheel.
c. Detach two wires from the terminal strip above
the map light. Note the connection and mark for reference when replacing the wires.
d. To install the control wheel map light reverse
this procedure.
e. For replacement of defective lamps, remove
two screws holding map light cover in place and unplug rheostat to remove cover.
f. Unsnap lamp sockets and replace lamps.
g. To reassemble, reverse this procedure.

CONTROL WHEEL MAP LIGHT.

17-93. DESCRIPTION. As optional equipment, a
white, dimmable map light may be installed on the
underside of the pilot's control wheel. On 1969
models, a solid-state dimming circuit along with a
miniature dimming control was used. On 1970
thru 1971 models, a new type of optional map light
has been installed on the underside of the pilot's
control wheel. The new map light assembly consists of a rectangle shaped housing containing two
small lamps and a small rheostat. On both type of
installations, the dimming control extends just below the edge of the control wheel map light housing
for convenient thumb or finger operation. For dimming the control should be rotated clockwise. Beginning with 1972 models the control wheel map
light is internally mounted in the control wheel. Thru
1974 models a rheostat switch located on the right
hand forward side of the wheel controls the light, Beg
Beginning with 1975 models the rheostat switch is
located on the lower right hand side of the control
wheel.
17-94. REMOVAL AND INSTALLATION (THRU U
206-1444) (Refer to Figure 17-15.)
a. Rotate the control wheel 90° to the left to gain
access to the underside of the control wheel.
b. Remove four screws at the corner of the etched
circuit board assembly.

17-38

It is recommended that the board be replaced
as an assembly if the lamps should become
defective. If personnel familiar with etched
circuit board repair work are available, emergency repairs of the map light assembly may
be made by soldering leads to #330 lamps
and then soldering the lamps to the board in
place of those provided. The lamps should be
secured in place with a spot of epoxy cement
after soldering.

Change 2

17-96. REMOVAL AND INSTALLATION. (AIRCRAFT
SERIAL U20601701 THRU U20601757).
a. Disconnect electrical cable connector of aft side
of control wheel.
b. Remove screws securing control wheel back
plate to control wheel tube adapter.
c. Remove screws securing plate to control wheel.
d. Disconnect socket from map light lamp and reflector unit.
e. Remove lamp and reflector unit.
NOTE
Lamp and reflector unit are bonded to control wheel.
CAUTION
Care must be taken in removing excess
bonding material, (do not hammer on
control wheel) as control wheel could be
damaged.
Using
f.
Conley Weld C and C2 or Hysol 5095 and
3673, bond new lamp and reflector unit.
g. To reassemble, reverse this procedure.

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

Tinnerman Nut
Grommet
Screw
Reflector
Socket
Bulb
Inspection Plate
Doubler
Lens
Spacer
Nutplate

1970 MODELS & ON

Detail
Figure 17-14.

A

Courtesy Light Installation
17 -39

2

THRU 1969 MODELS ONLY
NOTE
The "NAV LIGHTS" switch must be turned
on in order to operate the control wheel
map light.

1.
2.
3.

Screw
Lamp
Dimming Control

Figure 17-15.

4.
5.
6.

Map Light Housing
Transistor
Circuit Board

17-40

Change 2

Resistor
Terminal Board
Control Wheel

Control Wheel Map Light Installation (Sheet 1 of 4)

17-97. REMOVAL AND INSTALLATION. (BEGINNING WITH AIRCRAFT SERIAL U20601758 AND ALL
SERVICE PARTS BEGINNING WITH U20601701). To
remove, push upward on the lamp and turn. The lamp
and reflector is replaced as a unit.

SHOP NOTES:

7.
8.
9.

17-98.

COMPASS AND RADIO DIAL LIGHTS.

17-99. DESCRIPTION. The compass and radio dial
lights are contained within the individual units. The

NOTE
The "NAV LIGHTS" switch must be turned
on in order to operate the control wheel
map light.

3

Block
Cable

4

Terminal
1.
Nut 2.
Spectrastrip
3.

/14
12

Detail A

4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.

Sta-Strap
Screw
Control Wheel
Housing
Socket (Lamp)
Socket (Rheostat)
Plug Button
Lamp
Lens
Cover
Rheostat

1970 AND 1971 MODELS

Figure 17-15.

Control Wheel Map Light Installation (Sheet 2 of 4)
Change 2

17-41

NOTE

*

12
14

The "NAV LIGHTS" switch must be turned
on in order to operate the control wheel
map light.

/

1972 MODELS

15

1

4

1973 THRU 1974 MODELS

1.
2.
3.
4.
5.
6.
7.

Tube
Cover
Adapter
Rubber Cover
Plate
Map Light Rheostat
Terminal Block
Figure 17-15.

17-42

Change 2

8.
9.
10.
11.
12.
13.
14.

Map Light Assembly
Control Wheel
Pad
Mike Switch
Plug
Insulator
Electric Trim Switch

15.
16.
17.
18.
19.
20.
21.

plug
Bracket
Cable
Connector
Socket
Bracket
Lamp

Control Wheel Map Light Installation (Sheet 3 of 4)

2

14

13

BEGINNING WITH 1975 MODELS

1. Control Tube Assembly
2. Cover

8. Pad
9. Mike Switch

3. Adapter

10. Plug

4. Connector
5. Plate
6. Map Light Rheostat
7. Control Wheel

11. Insulator
12. Map Light Assembly
13. Lamp
14. Knob (Map Light)

Figure 17-15.

Control Wheel Map Light Installation (Sheet 4 of 4)
Change 2

17-42A/(17-42B blank)

light intensity is controlled by the radio dial light
dimming rheostat mounted on the lower left side of
the instrument panel.
17-100.

ELECTRIC CLOCK.

17-101. DESCRIPTION. The electric clock is connected to the battery through a 1-ampere fuse mounted adjacent to the battery box. The clock has a sweep
second hand and is an electro-mechanical type which
rewinds approximately every one and one-half minutes.
17-102.

STALL WARNING SYSTEM.

17-103. DESCRIPTION. The stall warning circuit
is comprised of a warning horn and an actuating
switch. The switch is installed in the leading edge
of the left wing and is actuated by airflow over the
surface of the wing. The switch will close as a
stall condition is approached, actuating the warning
horn which is mounted on the glove box. The stall
warning unit should actuate the stall warning horn
approximately five to ten miles per hour above the
aircraft stall speed. Install the lip of the warning
unit approximately one-sixteenth of an inch below
the centerline of the wing skin cutout. Test fly
the aircraft to determine if the unit actuates the

2

B

Detail

/

1.
2.
3.
4.
5.
6.

A
Wing Skin
Actuator
Tinnerman Nut
Screw
Map Compartment
Stall Warning Horn

3

Detail

B

*THRU

1971 MODELS

*BEGINNING WITH 1972 MODELS

Figure 17-16.

Stall Warning. Actuator and Horn Installation
Change 1

17-43

warning horn at the desired speed. If the unit actuates the warning horn at a speed in excess of ten
miles per hour above stall speed, loosen the mounting screws and move the unit down. If the unit actuates the horn five miles per hour below stall speed,
loosen the mounting screws and move the unit up.
17-104.

ice formations on the pitot tube and stall warning
actuator switch. The heaters are integrally mounted
in the pitot tube and the stall warning actuator switch.
Both heaters are operated by the pitot heat switch.
17-106. REMOVAL AND INSTALLATION OF PITOT
HEATER. Refer to Figure 17-17 for removal and
installation.

PITOT AND STALL WARNING HEATERS.

17-105. DESCRIPTION. Electrical heater units are
incorporated in some pitot tubes and stall warning
switch units. The heaters offset the possibility of

1.
2.
3.

17-107.

CIGAR LIGHTER.

17-108.

DESCRIPTION.

Electrical Leads
Pitot Tube
Heating Element

A special circuit breaker is

DetailA

Figure 17-17.

Pitot Heater Installation

1.
2.
3.
4
5.
6.

4

2
1

\
\

S

\

A\

\ T \
\7.

8.
S _><
99.

%J
10.

Figure 17-18.
17-44

Cigar Lighter Installation

Knob
Element
Socket
Panel
Shell
Circuit Breaker
Probe
Nut
Lockwasher
Power Wire

8

contained in a small cylinder screwed directly on
the back of the cigar lighter socket. The circuit
breaker is a bi-metallic type and is resettable. To
reset a breaker, make sure that the master switch
is off, then insert a small diameter pin (end of a
paper clip works) into the hole in the phenolic back
plate of the breaker and apply pressure. A small
click will be heard when the breaker resets,
17-109. REMOVAL AND INSTALLATION (Refer to
Figure 17-18.)
a. Ensure that the master switch is "OFF."
b. Remove cigar lighter element,
c. Disconnect wire on back of lighter,
d. Remove shell that screws on socket back of
panel.
e. The socket will then be free for removal,
f. To install a cigar lighter, reverse this procedure.
17-110.

SKYDIVING KIT.

17-111. DESCRIPTION. The kit consists of a
spoiler, sky diver steering switch, and a steering
signal light console. The spoiler is installed on the
door hinges of the removed front cargo door to mini-

2.
3. Screw
4.
5. Switch
6.

mize the strong air tlow buffeting within the cabin
when cargo doors are removed. The rocker-type
steering switch is mounted inside the cabin on the
upper sill of the cargo door opening and is used by
the sky diver to signal the pilot of his desired flight
path over the drop zone. A steering signal light
console, with red and green lights controlled by
operation of the steering switch, is mounted on top
of the instrument panel. Illumination of the red light
indicates to the pilot that the diver desires that the
aircraft be steered left; conversely, a green light
shows that the pilot is to steer right. Removal of
the cargo doors necessitates the installation of a
depressor plate over the wing flap circuit interrupt
switch to permit flap operation with doors removed.
(Under normal operations with the cargo door installed the switch prevents flap operation whenever
the front cargo door is open to prevent accidental
damage to the door or wing flap if the flaps are
lowered.)
17-112. REMOVAL AND INSTALLATION. For removal and installation of skydiving kit, refer to Figure 17-19. Refer to wing flap wiring diagrams in the
Wiring Section of this manual for wiring associated
with the flap circuit interrupt switch.

1 Spoiler Assembly
Angle - Flap Switch

9.
10.
11.
12.
13.
14.
15.

Pin
Bracket
Cover
Light Assembly - Right
Nutplate
Grommet
Bracket Assembly
Bulb
Spoiler Assembly

Figure 17-19.

DetailC

Sky Diving Components Equipment Installation
17-45

17-113. EMERGENCY LOCATOR TRANSMITTER.
17-114.

DESCRIPTION. The ELT is a self-contained,

solid state unit, having its own power supply, with an
externally mounted antenna. The C589510-0209 transmitter is designed to transmit simultaneously on dual
emergency frequencies of 121.5 and 243.0 Megahertz.
The C589510-0211 transmitter used for Canadian
registry, operates on 121.5 only. The unit is mounted in the tailcone, aft of the baggage curtain on the
right hand side. The transmitters are designed to
provide a broadcast tone that is audio modulated in a
swept manner over the range of 1600 to 300 Hz in a
distinct, easily recognizable distress signal for receptlon by serch and rescue personnel and others
monitoring the emergency frequencies. Power is
supplied to the transmitter by a battery-pack which
has the service life of the batteries placarded on the
batteries and also on the outside end of the transmitter. ELT's thru early 1974 models, were equipped
with a battery-pack containing six magnesium "D"
size dry cell batteries wired in series. (See figure
17-20) Mid 1974 thru early 1975, ELT's are equipped
with a battery-pack containing four "in-line" lithium
"D" batteries wired in series. Early 1975 and on
ELT's are equipped with a battery-pack containing
four lithium "D" size batteries which are stacked in
two's (See figure 17-22). The ELT exhibits line of
sight transmission characteristics which correspond
approximately to 100 miles at a search altitude of
10, 000 feet. When battery inspection and replacement
schedules are adhered to, the transmitter will broadcast an emergency signal at rated power (75 MWminimum), for a continuous period of time as listed
in the following table.
table.
TRANSMITTER LIFE
TO 75 MILLIWATTS OUTPUT

CAUTION
Do not leave the emergency locator transmitter

in the ON position longer than 5 seconds or
you may activate downed aircraft procedures
by C. A. P., D. O.T. or F. A. A. personnel.

WARNING
Magnesium (6-cell) battery-packs (excluding
4 cell lithium battery-packs) after prolonged
continuous use (1 hour) in a sealed environment give off explosive gas. If your ELT
has operated for this time period or longer,
as a precautionary measure, loosen the
ELT cover screws, lift the cover to break
air tight seal and let stand for 15 minutes
before tightening screws. Keep sparks,
flames and lighted cigarettes away from
battery-pack.
NOTE
After relatively short periods of inactivation,
the magnesium (6-cell) battery-pack develops
a coating over its anode which drastically
reduces self discharge and thereby gives
the cell an extremely long storage life.
This coating will exhibit a high resistance
to the flow of electric current when the
battery is first switched on. After a short
while (less than 15 seconds), the battery
current will completely dissolve this coating
and enable the battery to operate normally.
If this coating is present when your ELT is
activated, there may be a few seconds delay
before the transmitter reaches full power.
17-116. CHECKOUT INTERVAL:
100 HOURS.

Temperature

6 Cell
Magnesium
Battery Pack

4 Cell
Lithium
Battery Pack

+130°F

89 hrs

115 hrs

t 70°F

95 hrs

a. Turn aircraft master switch ON.
b. Turn aircraft transceiver ON and set frequency
on receiver to 121.5 MHz.
c. Remove the ELT's antenna cable from the ELT

- 4*F
- 40°F

115 hrs

49 hrs
23 hrs

unit.

95 hrs
70 hrs

Battery-packs have a normal shelf life of five to ten
(5-10) years and must be replaced at 1/2 of normal
shelf life in accordance with TSO-C91. Cessna
specifies 3 years replacement of magnesium (6-cell)
battery-packs and 5 years replacement of lithium
(4-cell) battery packs.
17-115. OPERATION. A three position switch on the
forward end of the unit controls operation. Placing
the switch in the ON position will energize the unit
to start transmitting emergency signals. In the OFF
position, the unit is inoperative. Placing the switch
in the ARM position will set the unit to start transmitting emergency signals only after the unit has
received a 5g (tolerances are +2g and -Og) impact
force, for a duration of 11-16 milliseconds.
17-46

Change 3

d. Place the ELT's function selector switch in the
ON position for 5 seconds or less. Immediately replace the ELT function selector switch in the ARM
position after testing ELT.
e. Test should be conducted only within the time
period made up of the first five minutes after any
hour.
CAUTION
Tests with the antenna connected should be
approved and confirmed by the nearest control
tower.
NOTE
Without its antenna connected, the ELT will
produce sufficient signal to reach your receiver,
yet it will not disturb other communications
or damage output circuitry.

A

PLACARD LOCATED ON UPPER R. H.
CORNER OF BAGGAGE CURTAIN
1

9

2

<

r10 *
1. Tailcone Skin

/

DetailB

2. Bracket
3. Transmitter
4. Battery Pack

A.

( Refer to paragraph 17-118. )

6. Cover

7 Connector
8. Arm Switch

K

9. Co-axial Cable

18

10. Sta-strap
11.

Antenna

15.

Metal Strap

.
I9

1

,

16. Suppressor

I
^^^'/ ;/

22
/

A

Detail
18

9'
s_

-3 .

3t, T<^^,~~
ransm<~i-^i~
>
1

/

>

e

.ll~k^^'-^s^^^~~

6~

Detail

.~

-

~
~~

r

^

7

1^

'

--'

/^{NOTE

Metal Strap (15) must be positioned so that
latch is on top of transmitter as installed
in the aircraft and not across transmitter

cover.

A

17. Placard
18. Fabric Fastener - hook
19. Fabric Fastener - Pile

Figure 17-20.

Emergency Locator Transmitter Installation
Change 3

17-47

NOTE
After accumulated test or operation time
equals 1 hour, battery-pack replacement
is required.
f. Check calendar date for replacement of batterypack. This date is supplied on a sticker attached to
the outside of the ELT case and to each battery.
17-117. REMOVAL AND INSTALLATION OF TRANSMITTER. (Refer to figure 17-20. )
a. Remove the baggage curtain to gain access
to the transmitter and antenna.
b. Disconnect co-axial cable from end of transmitter.
c. Depending upon the particular installation, either
cut four sta-straps and remove transmitter or cut
sta-strap securing antenna cable and unlatch metal
strap to remove transmitter.
NOTE
Transmitter is also attached to the mounting
bracket by velcro strips; pull transmitter to
free from mounting bracket and velcro.

a. Disconnect co-axial cable from base of antenna.
b. Remove the nut and lockwasher attaching the
antenna base ot the fuselage and the antenna will be
free for removal
c. To reinstall the antenna, reverse the preceding
steps.
NOTE
Upon reinstallation of antenna, cement
rubber boot (14) using RTV102, General
Electric Co. or equivalent, to antenna
whip only; do not apply adhesive to fuselage skin or damage to paint may result.
CAUTION
In-service 6 cell magnesium battery-pack
powered ELT's require the installation of a
static electricity suppressor in the antenna
cable to prevent the possibility of damage to
the case of the ELT. Refer to Cessna Avionics Service Letter AV74-16 and figure 17-20.
17-119. REMOVAL AND INSTALLATION OF MAGNESIUM SIX (6) CELL BATTERY-PACK. (Refer to
figure 17-21.)
NOTE

NOTE
To replace velcro strips, clean surface thoroughly with clean cloth saturated in one of the
following solvents: Trichloric thylene, Allphatic Napthas, Methyl Ethyl Ketone or Enmar 6094 Lacquer Thinner. Cloth should be
folded each time the surface is wiped to present a clean area and avoid redepositing of
grease. Wipe surface immediately with clean
dry cloth, do not allow solvent to dry on surface. Apply Velcro #40 adhesive to each surface in a thin even coat and allow to dry until
quite tacky, but no longer transfers to the
finger when touched (usually between 5 and
30 minutes). Porous surfaces may require
two coats. Place the two surfaces in contact
and press firmly together to insure intimate
contact. Allow 24 hours for complete cure.

On aircraft incorporating Cessna ELT's
manufactured by Leigh (Shark 7 series),
when replacing battery-pack refer to
Cessna Avionics Service Letter AV75-5,
dated July 3, 1975.

Since replacement 6 cell magnesium batterypacks are no loger available, when inservice units require replacement, use the
4cell lithium battery-pack. Refer to paragraph
17-120.
TRANSMITTER
C589510-0102

e. To reinstall transmitter, reverse preceding
steps.
NOTE
An installation tool is required to properly
secure sta-straps on units installed with
sta-straps. This tool may be purchased
locally or ordered from the Pandiut Corporation, Tinley Park, 111., part number
GS-2B (Conforms to MS90387-1).

BATTERY-PACK
ELECTRICAL
CONNECTOR

C589510-0105
(6 Cell Magnesium)

CAUTION
Ensure that the direction of flight arrows
(placarded on the transmitter) are pointing
towards the nose of the aircraft.
17-118. REMOVAL AND INSTALLATION OF
ANTENNA. (Refer to figure 17-20.)
17-48

Change 3

Figure 17-21. Magnesium 6 Cell
Battery-Pack Installation
17-120. REMOVAL AND INSTALLATION OF LITHIUM
FOUR (4) CELL BATTERY-PACK. (Refer to figure
17-22.)

NOTE
On aircraft incorporaring Cessna ELT's
manufactured by Leigh (Shark 7 series),
when replacing battery-pack refer to
Cessna Avionics Service Letter AV75-5,
dated July 3. 1975.

CAUTION
Be sure to enter the new battery-pack expiration date in the aircraft records. It is also
recommended this date be placed in your ELT
Owner's Manual for quick reference.

NOTE
Transmitters equipped with the 4 cell batterypack can only be replaced with another 4 cell
battery-pack.

TRANSMITTER
C589510-0202
(4

BATTERY PACK
C589510-0205
Cell Lithium)

a. After the transmitter has been removed from
aircraft in accordance with para. 17-117, place the
transmitter switch in the OFF position.
b. Remove the nine screws attaching the cover to
the case and then remove the cover to gain access to
the battery-pack.
NOTE
Retain the rubber "0" ring gasket, rubber
washers and screws for reinstallation.

ELECTRICAL
CONNECTOR

MELT
ADHESIVE
3M (PN 3738)

c. Disconnect the battery-pack electrical connector
and remove battery-pack.
d. Place new battery-pack in the transmitter with
four batteries as shown in the case in figure 17-22.
e. Connect the electrical connector as shown in figure 17-22.
NOTE
Before installing the new 4 cell batterypack, check to ensure that its voltage is
11. 2 volts or greater.
CAUTION
If it is desireable to replace adhesive material on the 4 cell battery-pack, use only 3M
Jet Melt Adhesive 3738. Do not use other
adhesive materials since other materials
may corrode the printed circuit board assembly.
1. Replace the transmitter cover by positioning the
rubber "0" ring gasket. if installed, on the cover
and pressing the cover and case together. Attach
cover with nine screws and rubber washers.
g. Remove the old battery-pack placard from the
end of transmitter and replace with new battery-pack
placard supplied with the new battery-pack.

TRANSMIITTERBATTERY PACK
C589510-0209
C589510-0210

Figure 17-22. Lithium 4 Cell
Battery Pack Installations

17-121. TROUBLE SHOOTING. Should your Emergency Locating Transmitter fail the 100 Hours performance checks, it is possible to a limited degree
to isolate the fault to a particular area of the equipment. In performing the following trouble shooting
procedures to test peak effective radiated power.
you will be able to determine if battery replacement
is necessary or if your unit should be returned to
your dealer for repair.

SHOP NOTES:

Change 3

17-49

TROUBLE
POWER LOW

PROBABLE CAUSE

REMEDY

Low battery voltage.

1. Set toggle switch to off.
2. Remove plastic plug from the remote jack
and by means of a Switchcraft #750 jackplug.
connect a Simpson 260 model voltmeter and
measure voltage. If the battery-pack voltage
on the 6-cell magnesium battery pack transmitter is 10.8 volts or less, and on the 4-cell
lithium battery pack transmitters is 11.2 volts
or less, the battery pack is below specification.

Faulty transmitter.

3. If the battery-pack voltage meets the
specifications in step 2, the battery-pack
is O. K. If the battery is O. K., check the
transmitter as follows:
a. Remove the voltmeter.
b. By means of a switchcraft 750 jackplug
and 3 inch maximum long leads, connect a
Simpson Model 1223 ammeter to the jack.
c. Set the toggle switch to ON and observe
the ammeter current drain. If the currentdrain is in the 85-100 ma range. the
transmitter or the co-axial cable is faulty.

Faulty co-axial
antenna cable.

4. Check co-axial antenna cable for high
resistance joints. If this is found to be
the case, the cable should be replaced.

*This test should be carried out with the co-axial cable provided with your unit.

SHOP NOTES:

17-50

Change 3

ELECTRICAL LOAD ANALYSIS CHART
24 VOLT ALL MODELS
AMPS REQD

STANDARD EQUIPMENT (RUNNING

LOAD)

1971

1972

1973

1974

1975

1976

0.6
t
0.2
0.4
7.0

.41
t
.039
.12
6.0

.41
t
.039
.12
6.0

.41
t
0.039
0.12
4.0

.41
t
0.039
0.12
4.0

.41
0.039
0.12
4.0

.03
0.2
1.0
.04
2.0
0.4

.03
0.2
1.0
.04
2.0
.28

.03
0.2
1.0
.04
2.0
.28

0.02
0.16
1.14
0.04
2.0
0.3

0.02
0.16
1.14
0.04
2.0
0. 3

0.02
0.16
1.14
0.04
2.0
0.3

Heated-Pitot ..
.
5.8
Strobe Lights . . . . .. . ..
. ...
. ....
4.0
..
0.03
..
Carburetor Air Temp . . .......
.
Cessna 200A Navomatic (Type AF-295A) ..........
Cessna 200A Navomatic (Type AF-295B) ...........
1.6
....
Cessna 300 ADF (Type R-521B)
..
.........
Cessna 300 ADF (Type R-546A) ...
..........
Cessna 300 ADF (Type R-546E) .
02
Cessna 300 Marker Beacon (Type R-502B) ..........
Cessna 300 Nav/Com (90 Channel-Type RT-517R) ..
4.5
4.5
Cessna 300 Nav/Com (360 Channel-Type RT-540A) . .....
Cessna 300 Nav/Com (100 Channel-Type RT-508A). .....Cessna 300 Nav/Com (360 Channel-Type RT-308C). .....
Cessna 300 Nav/Com (360 Channel-Type RT-528A). .....
Cessna 300 Nav/Com (360 Channel-Type RT-528E). ....
Cessna 300 Nav/Com (360 Channel-Type RT-328A). ....
Cessna 300 Nav/Com (360 Channel-Type RT-328C). .....
...
Cessna 300 Nav/Com (720 Channel-Type RT-328D) .
Cessna 300 Transceiver (Type RT-524A) .....
.....
Cessna 300 HF Transceiver (Type PT-10A) .........
.0.7
Cessna 300 Transponder (Type KT-75R) .
....
Cessna 300 Transponder (Type KT-76 & KT-78) .
Cessna 300 Transponder (Type RT-359A) ..........
1.8
Cessna 300 Navomatic (Type AF-512C) .
....
..
Cessna 300 Navomatic (Type AF-512D) .
.
Cessna 300 Navomatic (Type AF-394A) ........
Cessna 300A Navomatic (Type AF-395A) ..........3.0
Cessna 300 DME (Type KN-60B) .............
......Cessna 300 DME (Type KN-60C) .
.1.8
...
Cessna 400 ADF (Type R-324A)
.
.
.. Cessna 400 ADF (Type R-346A)
.
Cessna 400 ADF (Type R-446A) ...........
0.4
Cessna 400 Glideslope (Type R-.543B). ...........
...........
Cessna 400 Glideslope (Type R-443A).
Cessna 400 Glideslope (Type R-443B). ...........
Cessna 400 Nav/Com (Type RT-522A). ...........
3.0
. Cessna 400 Nav/Com (Type RT-422A). ..........
2.2
....
Cessna 400 Transceiver (Type RT-532A) .
Cessna 400 Transceiver (Type RT-432A) ..........
1.5
.
........
Cessna 400 Transponder (Type RT-506A) .
Cessna 400 Transponder (Type RT-459A) ..........
Cessna 400 Nav-O-Matic (Type AF-520C) ..........
1.2
Cessna 400 Nav-O-Matic (Type AF-420A) ..........
-

5.8
4.0
0.03
-

5.8
4.0
0.03
--

5.8
4.0
0.03
1.5
-

5.8
4.0
0.03

5.8
4.0
0.03

-

-

--

1.0
1.0
.02

1.0
1.0
02

1.0
1.0
0.02

1.0
1.0
0.002

1.5
1.0
1.0
0.02

1.5

1.5

1.5

-

-

Battery Contactor ..
.
...
Clock .
.
.
....................
Cylinder Head Temperature Indicator
........
Fuel Quantity Indicators
.
.
..............
Flashing Beacon .
.
.................
Instrument Lights
a. Electroluminescent Panel ........
...
b.
Cluster ............
.......
c.
Console* .
. . . . . .....
. . . . . . . ..
d.
Compass . . ..
..
..
. ..
...
. ..
...
Position Lights
.
. . . . ..
...........
Turn Coordinator
. . ..
.
.
..
. ....
...
. . .

OPTIONAL EQUIPMENT (RUNNING

.

LOAD)

-

--

-

-

1.9

1.9
1.9

-

1.9
- 1.9
1.9
--

-

1.9

1.9 1.9
1.5

-

-

-

1.5
2.1

1.0

1.0

1.5
2.1
-

-

-

-

1.3
-

-

-

1.0

1.0

-

-

-

-

1.8
-

1.75
-

-

-

-

2.0

2.0

-

-

-

3.0

3.0

2.4

-

-

-

-

-

-

1.0
0.4
3.0
2.2

1.0
0.4
0.4
3.0
2.5

1.0
0.4

--

0.4

-

-

0.32
3.0
1.7

0.32
3.0
-

-

-

1.7
1.5
-

-

-

-

1.0

1.0

1.0

-

-

-

1.2

1.2

1.2

1.0
0.7
1.3
-

1.5
- -1.2
-

1.8
-

1.4

Change 3

-

1.0

-

1.0
0.4
0.32
3.0
-

1.2

17-51

ELECTRICAL LOAD ANALYSIS CHART (CONT.)
24 VOLT ALL MODELS

AMPS REQD

OPTIONAL EQUIPMENT (RUNNING LOAD)
(CONT.)
Cessna 400 Area Nav (Type RN-478A). ..........
Cessna 400 DME (Type R-476A) .............
Bendix MKR BCN RCVR (Type GM-247A) .........
King KN-65 DME .
........
...........
Sunair SSB Transceiver (Type ASB-125) .
........
Narco Mark 12B Nav/Com with VOA-40 or VOA-50
Narco UGR-2 Glideslope Receiver ............
King KN-60C DME ...................
Pantronics PT-IOA HF Transceiver ............

.

....
.

1971

1972

1973

2.5
4.6
.23
-

2.5
-

2.5
-

2.5
-

3.0
7.0
8.5
3.57
1.0
.25

3.0
7.0
8.5
3.57
1.0
.25

3.0
7.0
8.5
3.57
1.0
.25

1.2
.04

1.2
.04

1.2
.04

1974
-

1975

1976

1.4
2.5
2.4
1.5

0. 5
2.5
.100
1.4
2.5

3.0
7.0
8.55
3.57
1.0
.28

3.0
7.0

3.0
7.0

3.57
1.0
.28

3.57
1.0
.28

1.65
.04

1.65
.04

1.65
.04

2.4
1.5

ITEMS NOT CONSIDERED AS PART OF
RUNNING LOAD.
Auxiliary Fuel Pump ..................
Cigarette Lighter. ..
.......
Flap Motor ......................
Landing Lights (Each) .................
Oil Dilution System . ............
Stall Warning Horn .....

......

. . .
.

.

.

Wing Courtesy Lights and Cabin Lights . ......
Sky Diving Lights . . .............

...

. .

.

8.5

*Console lights not used with post lights.
Only one or the other may be used at one time.
±Negligible

12 VOLT ALL MODELS

STANDARD EQUIPMENT
(RUNNING

AMPS REOD

LOAD)

1969

1970

1971

1972

1973

1974

1975

.6
t

0.6
t

0.6
t

0. 6
t

0.6
t

0.6
t

0.6
t

0.2
0.4
7.0

0.2
0.4
7.0

2
0.4
7.0

0.2
0.4
7.0

0.2
0.4
7.0

0.2
0.4
7.0

0.2
0.4
7.0

0.5
0.3
2.0
0.1
5.6
0. 8

0. 5
0.3
2.0
0.1
5.6
0.8

0. 5
0.3
2.0
0.1
5.6
0.8

0. 5
0.3
2.0
0.1
5.6
0.8

0. 5
0.3
2.0
0.1
5.6
0.8

0.4
0.32
2.08
0.8
5.6
0.8

0.4
0.32
2.08
0.8
5.6
0.8

10.0

10.0

10.0

Battery Contactor ................
Clock
. . . . . . . . . . . . . . . . . . . . .
Cylinder Head Temperature Indicator ....
Fuel Quantity Indicators
.............
Flashing Beacon .................
Instrument Lights ........
a.
Electroluminescent Panel .........
b. Cluster .................
c. Console* ................
d.
Compass .
.
........
Position Lights ..................
Turn Coordinator .................

.

.

OPTIONAL EQUIPMENT
(RUNNING LOAD)
Heated-Pitot, Stall Warning Heater .........

Strobe Lights

..................

.-

Carburetor Air Temp ..............
Cessna 200A Navomatic Autopilot (Type AF-295A) . .
Cessna 200A Navomatic Autopilot (Type AF-295B) .
Cessna 300 ADF (Type R-521B) ..........

17-52

Change 3

0. 03

0.03

10.

1 . 0 0.

1.6

0.0

4.0

4.0

2.0

2.0

0. 03

0. 03

0. 03

0.03
2.0
--

0. 03
2.0

-1.6

0 0.0

4.0

ELECTRICAL LOAD ANALYSIS CHART (CONT.)
12 VOLT ALL MODELS

OPTIONAL EQUIPMENT

1969

LOAD) (CONT.)
Cessna 300 ADF (Type R-546A) ..........
.
Cessna 300 ADF (Type R-546E) .
...
. ..
Cessna 300 Marker Beacon (Type R-502B) .....
Cessna 300 Nay/Com (90 Channel-Type RT-517R) .
Cessna 300 Nav/Com (360 Channel-Type RT-540A).
Cessna 300 Nav/Com (100 Channel-Type RT-508A).
Cessna 300 Nav/Com (360 Channel-Type RT-308C).
Cessna 300 Nav/Conm (360 Channel-Type RT-528A). .Cessna 300 Nav/Corn (360 Channel-Type RT-528E).
Cessna 300 Nav/Cornm (360 Channel-Type RT-328A).
Cessna 300 Nav/Com (360 Channel-Type RT-328C).
Cessna 300 Nav/Com (720 Channel-Type RT-328D)
Cessna 300 Transceiver (Type RT-524A) .....
.
Cessna 300 HF Transceiver (Type PT-10A) .
..
.
Cessna 300 Transponder (Type KT-75R) ...
. . .
Cessna 300 Transponder (Type KT-76 & KT-78) . .
Cessna 300 Transponder (Type RT-359A) ...
. . .
Cessna 300 Navumatic (Type AF-512C) .......
Cessna 300 Navomiatic (Type AF-512D) .......Cessna 300 Navomatic (Type AF-394A) ......
. . . .
Cessna 300A Navomatic (Type AF-395A) ..
...
.
....
Cessna 300 DME (Type KN-60B)
Cessna 300 DME (Type KN-60C) . . .
.
. .
. . . . .
Cessna 400 ADF (Type R-324A) ..
... .
.
Cessna 400 ADF IType R-346A) . ..
.........Cessna 400 ADF (Type R-446A) .
...
Cessna 400 Glideslope (Type R-543B)
Cessna 400 Glideeslope (Type R-443A) .
...
Cessna 400 Ghldeslupe (Type R-443B) ........
. . .. ..
Cessna 400 Nay/Curn (Type RT-522A)
Cessna 400 Nav.Cuni (Tvpe RT-422A). .......--Cessna 400 Transceiver (Type RT-532A) ......
Cessna 400 Transceiver (Type HT-432A . . . . .
Cessna 400 Transponder (Type RT-506A) .....
.Cessna 400 Transponder (Type RT-459A) ...
Cessna 400 Nav-O-Matic (Type AF-520C) . .....
Cessna 400 Nav-O-Maric (Type AF-420A . .....
Sunair SSB Transceiver 'Tvpe ASB-125 . ......
...........
.
Flashinc Beacon
Kin- KN-60C DME .
.....
..-Kinm KN-65 DME . ...
pantronics PT-10A HF Transceiver
........Narcu Mark 12A Navy/Corn .............
Narcu Mark 12B Nayv/Com with VOA-40 or VOA-50. .
Narco UGR-2 Glideslope Receiver . . ......

ITEMS NOT CONSIDERED
OF RUNNING LOAD

.02
4.5
4.5
-

1970
- -

-

-

-

.02
4.5
4.5
---

.02
4.5
4.5
--

-

-

-

3.2
1.5
-3. 5

3.2
1.5

3.0
2.0

1.0
1.0
.02

-

-

1.9
1.9
--

--

1.0
1.0
.02

-

3.5
3.0
2.0

3.2
1.5
1.5
3.5
3.02.0

3.2
1.5
1.5
1.3
-

3.5
-

0.5

3.

3.0

1.5
3.0

.5 1.5
1
3.0
3.0
2.4
5.0
5.0
7.0
7.0

1 .5
3.0
2.4
5.0
7.0

--

-

3.0
-

-

7.0

4.6
4.6
.23

-

-

-

4.6
.23

4.6
.23

1.0
1.0
0.02

-

-

1.5

1.5

-

-

1.9
1--1.5
3.2
1.5

1.3
-

-

-

-

2.0

1.9
-

3.2
1.5

3.0

1.0

1.5
3.2

1.0
-

2.0
3.0
-

-

0. 5
--

2.0

1.0
1.0
0.02

-

3.0
-

0.5

0.5

1915

2.0
-

1.0

-

1.9
1.9
1,-1.9
1.9
-

\I T4

1.0

1.0

-

-

-

0.5
0.4

0.5

0. 5

-

-

0.4
.0
2.5

0.4
3.0

-

.0
2.5

-

-

1.4
3.0
-

-

1.4
1.0

1.0

1.2
5.0
7.0

1.2
5.0
7.0

1.2
5.0

-

-

2.8

0........
2.R
1.5

--

AS PART

.....
......
Auxiliary Fuel Pump . ..
.
. . . . . .....
..
Cigarette Lighter
Flap Motor ..
. ...
.......
.
.

Landing Lights ...............
Oil Dilution System . .

AMPS REQD
12
4112

(RUNNING

........

.

3.0
10.0
15.0

3.0
10.0
15.0

3.0
10.0
15.0

3.0
10.0
15.0

3.0
10.0
15.0

3.0
10.0
15.0

3.0
10.0
15.0

15.6

15.6
1.0

15.6
1.0

15.6
1.0

15.6
1.0

15.6
1.0

15.6
1.0

.1.0

Stall Warning Horn ............

....

0.25

0.25

0.25

0.25

0.25

0.25

0.25

Wing Courtesy Lights and Cabin Lights .

.

3.3

3.3

3.3

3.3

3.3

3. 3

3.3

0.1

0.1

0.1

0.1

0.

0.

0.

Sky Diving Lights.

.

. .....

. ..

*Console lights not used with post lights.
Only one or the other may be used at one time
tNegligible

1
Change 2

17-53/(17-54 blank)

SECTION 18
STRUCTURAL REPAIR
TABLE OF CONTENTS

Page

STRUCTURAL REPAIR ...........
Repair Criteria ................
...
Equipment and Tools .......
Control Balancing Fixtures ......
Support Stands .........
Fuselage Repair Jig .........
Wing Jig ..........
Wing Twist and Horizontal Stabilizer
Angle-of-Incidence ..........
Repair Materials ..........
Wing .................
Skin.
...............
Negligible Damage .........
Repairable Damage ...........
Damage Necessitating Replacement of Parts ........
Stringers ..............
Negligible Damage .........
Repairable Damage .........
Damage Necessitating Replacement of Parts ........
Ribs ........
Negligible Damage .........
Repairable Damage .........
Damage Necessitating Replacement of Parts ........
Spars .................
Negligible Damage .........
Repairable Damage .........
Damage Necessitating Replacement of Parts ........
Leading Edge ............
Negligible Damage .........
Repairable Damage .........
Damage Necessitating Replacement of Parts ........
Bonded Leading Edge Repair
.....
Negligible Damage ........

18-1
18-1
18-1
18-1
18-1
18-1
18-2

18-1.

..

18-2
. 18-2
18-2
18-2
18-2
18-2
18-2
18-2
18-2
18-2
18-2
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3
18-3

REPAIR CRITERIA.

18-2. Although this section outlines repair permissible on structure of the aircraft, the decision of
whether to repair or replace a major unit of structure
will be influenced by such factors as time and labor
available, and by a comparison of labor costs with
the price of replacement assemblies. Past experience indicates that replacement, in many cases, is
less costly than major repair. Certainly, when the
aircraft must be restored to its airworthy condition
in a limited length of time, replacement is preferable;
18-3. Restoration of a damaged aircraft to its original design strength, shape and alignment involves
careful evaluation of the damage, followed by exacting workmanship in performing the repairs. This
section suggest the extent of structural repair practical on the aircraft and supplements Federal Aviation
Regulations, Part 43. Consult the factory when in
doubt about a repair not specifically mentioned here.

Repairable Damage .........
18-3
Ailerons ........
.
18-3
18-3
.
Negligible Damage ...
Repairable Damage .
........
18-3
Damage Necessitating Replacement of Parts ........
18-3
Flaps ....
...
..
18-3
Negligible Damage .......
. 18-3
Repairable Damage .........
18-3
Damage Necessitating Replacement of Parts ........
18-3
Elevators and Rudders .......
. 18-3
Negligible Damage .
.
........
18-3
Repairable Damage .
.
........
18-3
Damage Necessitating Replacement of Parts ..
.....
. 18-3
Fin and Stabilizer .
..........
18-4
Negligible Damage
..
....
.
18-4
Repairable Damage
.
.
....
.
18-4
Damage Necessitating Replacement of Parts .
.
.......
18-4
Fuselage .
..............
18-4
Negligible Damage .......
. 18-4
Repairable Damage .
.........
18-4
Damage Necessitating Re.
....
. 18-4
placement of Parts ..
18-4
Bulkheads ...............
18-4
Landing Gear Bulkheads ......
18-4
Repair after Hard Landings ......
18-4
Replacement of Hi-Shear Rivets .....
Nose Gear Wheel Well and Firewall ... .18-5
.
18-5
Baffles .............
. 18-5
Engine Cowling ...........
18-5
Repair of Cowling Skins ........
. 18-5
Repair of Reinforcement Angles ...
18-5
Repair of ABS Components .......
. .18-5
Repair of Glass-Fiber Constructed .
.
18-5
.......
Components .
Bonded Doors .......
...
. 18-5
Repairable Damage .
..
. ...
. 18-5
18-4.

EQUIPMENT AND TOOLS.

18-5. Equipment and tools for repair of structure
may be fabricated locally for all but major repair
jobs. For major repair of wings and fuselage,
special jigs, available from the factory are recommended. These jigs are precision equipment designed to ensure accurate alignment of these airframe components.
18-6. CONTROL BALANCING requires the use of
a fixture to determine the static balance moment of
the control surface assembly. Plans for, and the
use of, such a fixture are shown in figure 18-9.
18-7. SUPPORT STANDS shown in figure 18-1 are
used to hold a fuselage or wing when it is removed.
The stands may be manufactured locally of any suitable wood.
18-8. FUSELAGE REPAIR JIG. The fuselage jig,
which may be obtained from the factory, is a sturdy,
D2007C3-13 Temporary Change 1
Sheet 1 of 2
September 5/77

18-1

versatile fixture used to hold an entire fuselage and
to locate the firewall, wing and landing gear attachment points. The jig is ideal for assembling new
parts in repair of a badly damaged fuselage.

ings afford access to the aileron bellcranks, flap
bellcranks, electrical wiring, strut attaching fittings,
aileron control cable pulley and control cable disconnect points.

18-9. WING JIG. The wing jig, which may also be
obtained from the factory, serves as a holding fixture during extensive repair of a damaged wing. The
jig locates the root rib, leading edge, and tip rib of
the wing.

18-16.

18-10. WING TWIST AND STABILIZER ANGLE-OFINCIDENCE.Wing twist (washout) and horizontal stabilizer angle of incidence are shown below. Stabilizers do
not have twist. Wings have no twist from the root to
the lift strut station. All twist in the wing panel
occurs between this station and the tip rib. Refer to
figure 18-2 for wing twist measurement.
WING
Twist (Washout)

°

STABILIZER
Angle of Incidence

18-11.

3°
-3 ° 30'

REPAIR MATERIALS.

18-12. Thickness of material on which a repair is
to be made can easily be determined by measuring
with a micrometer. In general, material used in
Cessna aircraft covered in this manual is made from
2024 aluminum alloy, heat treated to a -T3, -T4, or
-T42 condition. If the type of material cannot be
readily determined, 2024-T3 may be used in making
repairs, since the strength of -T3 is greater than
-T4 or -T42 (-T4 and -T42 may be used interchangeably, but they may not be substituted for -T3). When
necessary to form a part with a smaller bend radius
than the standard cold bending radius for 2024-T4,
use 2024-0 and heat treat to 2024-T42 after forming.
The repair material used in making a repair must
equal the gage of the material being repaired unless ,
otherwise noted. It is often practical to cut repair
pieces from service parts listed in the Parts Catalogs.
A few components (empennage tips, for example) are
fabricated from thermo-formed plastic or glass fiber
constructed materials.
18-13.

WING.

18-14. The wing assemblies are of the semi-cantilever type employing semi-monocoque type of structure. Basically, the internal structure consists of
built-up front and rear spar assemblies, formed
sheet metal nose, intermediate, and trailing edge
ribs. Stressed skin, riveted to the rib and spar
structures, completes the wing structure.
18-15. ACCESS openings (hand holes with removable
cover plates) are located in the underside of the wing
between the wing root and tip section. These open18-2

D2007C3-13 Temporary Change 1
Sheet 2 of 2
September 5/77

WING SKIN.

18-17. NEGLIGIBLE DAMAGE. Any smooth dents
in the wing skin that are free from cracks, abrasions
and sharp corners, which are not stress wrinkles
and do not interfere with any internal structure or
mechanism, may be considered as negligible damage. In areas of low stress intensity, cracks, deep
scratches or deep, sharp dents, which after trimming or stop drilling can be enclosed by a two-inch
circle, can be considered negligible if the damaged
area is at least one diameter of the enclosing circle
away from all existing rivet lines and material edges.
Stop drilling is considered a temporary repair and a
permanent repair should be made as soon as practicable.
18-18. REPAIRABLE DAMAGE. Figure 18-3 outlines typical repairs to be employed in patching skin.
Before installing a patch, trim the damaged area to
form a rectangular pattern, leaving at least a onehalf inch radius at each corner, and deburr. The
sides of the hole should lie span-wise or chord-wise.
A circular patch may also be used. If the patch is in
an area where flush rivets are used, make a flush
patch type of repair; if in an area where flush rivets
are not used, make an overlapping type of repair.
Where optimum appearance and airflow are desired,
the flush patch may be used. Careful workmanship
will eliminate gaps at butt-joints; however, an epoxy
type filler may be used at such joints.
18-19. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If a skin is badly damaged, repair
should be made by replacing an entire skin panel,
from one structural member to the next. Repair
seams should be made to lie along existing structural
members and each seam should be made exactly the
same in regard to rivet size, spacing, and pattern
as the manufactured seams at the edges of the original sheet. If the manufactured seams are different,
the stronger should be copied. If the repair ends at
a structural member where no seam is used, enough
repair panel should be used to allow an extra row of
staggered rivets, with sufficient edge margin, to be
installed.
18-20.

WING STRINGERS.

18-21.
18-17.

NEGLIGIBLE DAMAGE.

Refer to paragraph

18-22. REPAIRABLE DAMAGE. Figure 18-4 outlines a typical wing stringer repair. Two such repairs may be used to splice a new section of stringer
material in position, without the filler material.
18-23. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If a stringer is so badly damaged that
more than one section must be spliced into it, replace the entire stringer.

18-24.

WING RIBS.

18-25.
18-17.

NEGLIGIBLE DAMAGE.

18-26. REPAIRABLE DAMAGE.
lines typical wing rib repairs.

Refer to paragraph
Figure 18-5 out-

18-27. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Leading edge and trailing edge ribs that
are extensively damaged should be replaced. However, due to the necessity of unfastening so much
skin in order to replace ribs, they should be repaired
if practicable. Center ribs, between the front and
rear spars should always be repaired if practicable.
18-28.

18-36.

AILERONS.

18-37.
18-17.

NEGLIGIBLE DAMAGE.

Refer to paragraph

18-38. REPAIRABLE DAMAGE. The repair shown
in figure 18-8 may be used to repair damage to aileron leading edge skins. Figure 18-3 may be used
as a guide to repair damage to flat surface between
corrugations, when damaged area includes corrugations refer to figure 18-11. It is recommended that
material used for repair be cut from spare parts of
the same guage and corrugation spacing. Refer to
figure 18-10 for balancing. If damage would require
a repair which could not be made between adjacent
ribs, refer to paragraph 18-39.

WING SPARS.

18-29. NEGLIGIBLE DAMAGE. Due to the stresses
which wing spars encounter, very little damage can
be considered negligible. All cracks, stress wrinkles,
deep scratches, and sharp dents must be repaired.
Smooth dents, light scratches, and abrasions may be
considered negligible.
18-30. REPAIRABLE DAMAGE. Figure 18-6 outlines typical spar repairs. It is often practical to
cut repair pieces from spare parts listed in Parts
Catalogs. Service Kits are available for certain
types of spar repairs.
18-31. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Damage so extensive that repair is not
feasible requires replacement of a complete wing
spar. Also refer to paragraph 18-2.
18-32.

WING LEADING EDGE.

18-33.
18-17.

NEGLIGIBLE DAMAGE.

Refer to paragraph
of

18-34. REPAIRABLE DAMAGE. A typical leading
edge skin repair is shown in figure 18-8. An epoxy
type filler may be used to fill gaps at butt joints. To
facilitate repair, extra access holes may be installed
in the locations noted in figure 18-7. If the damage
would require a repair which could not be made between adjacent ribs, refer to the following paragraph.
18-35. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. For extensive damage, complete leading edge skin panels should be replaced. To facilitate replacement, extra access holes may be installed
in the locations noted in figure 18-7.
18-35A.

BONDED LEADING EDGE REPAIR.

18-35B.
18-17.

NEGLIGIBLE DAMAGE.

Refer to paragraph
cludes

18-35C. REPAIRABLE DAMAGE. (Refer to figure
18-12.) Cut out damaged area, as shown, to the
edge of undamaged ribs. Using a corresponding
section from a new leading edge skin, overlap ribs
and secure to wing using rivet pattern as shown in
the figure.

18-39. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If the damage would require a repair
which could not be made between adjacent ribs, complete skin panels should be replaced. Ribs and spars
may be repaired, but replacement is generally preferable. Where extensive damage has occurred, replacement of the aileron assembly is recommended.
After repair and/or repainting, balance in accordance with figure 18-9.
18-40.

WING FLAPS.

18-41.
18-17.

NEGLIGIBLE DAMAGE.

Refer to paragraph

18-42. REPAIRABLE DAMAGE. Flap repairs should
be similar to aileron repairs discussed in paragraph
18-38. A flap leading edge repair is shown in figure
18-8.
18-43. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Flap repairs which require replacement
parts should be similar to aileron repairs discussed in paragraph 18-39.
18-44.

ELEVATORS AND RUDDERS.

18-45. NEGLIGIBLE DAMAGE. Refer to paragraph
18-17. The exception of negligible damage on the
elevator surfaces is the front spar, where a crack
appearing in the web at the hinge fittings or in the tip
rib which supports the overhanging balance weight is
not considered negligible. Cracks in the overhanging
tip rib, in the area at the front spar intersection with
the web of the rib, also cannot be considered negligible.
18-46. REPAIRABLE DAMAGE. Skin patches illustrated in figure 18-3 may be used to repair skin damage to the rudder, and between corrugations on the
elevator. For skin damage on the elevator which incorrugations, refer to figure 18-11. Following repair the elevator/rudder must be balanced.
Refer to figure 18-10 for balancing. If damage would
require a repair which could not be made between
adjacent ribs, refer to paragraph 18-47.
18-47. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If the damaged area would require a repair which could not be made between adjacent ribs,
Change 2

18-3

complete skin panels should be replaced. Ribs and
spars may be repaired, but replacement is generally
preferable. Where extensive damage has occurred,
replacement of the entire assembly is recommended.
Alter repair and/or repainting, balance in accordance with figure 18-9.
18-48.

FIN AND STABILIZER.

18-49.
18-17.

NEGLIGIBLE DAMAGE.

Refer to paragraph

18-50. REPAIRABLE DAMAGE. Skin patches shown
in figure 18-3 may be used to repair skin damage.
Access to the dorsal area of the fin may oe gained
by removing the horizontal closing rib at the bottom
of the fin. Access to the internal fin structure is
best gained by removing skin attaching rivets on one
side of the rear spar and ribs, and springing back
the skin. Access to the stabilizer structure may be
gained by removing skin attaching rivets on one side
of the rear spar and ribs, and springing back the
skin. If the damaged area would require a repair
which could not be made between adjacent ribs, or
a repair vould be located in an area with compound
curves, see the following paragraph.
18-51. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If the damaged area would require a repair which could not be made between adjacent ribs
or the repair would be located in an area with compound curves, complete skin panels should be replaced. Ribs and spars may be repaired, but replacement is generally preferable. Where damage is extensive, replacement of the entire assembly is recommended.
18-52.

FUSELAGE.

18-53. The fuselage is of semi-monocoque construction consisting of formed bulkheads, longitudinal
stringers, reinforcing channels and skin platings.
18-54. NEGLIGIBLE DAMAGE. Refer to paragraph
18- 17. Mild corrosion appearing upon alclad surfaces does not necessarily indicate incipient failure
of the base metal. However, corrosion of all types
should be carefully considered, and approved remedial action taken. Small cans appear in the skin
structure of all metal airplanes. It is strongly recommended, however, that wrinkles which appear to
have originated from other sources, or which do not
follow the general appearance of the remainder of the
skin panels, be thoroughly investigated. Except in
the landing gear bulkhead area, wrinkles occurring
over stringers which disappear when the rivet patHowtern is removed may be considered negligible.
ever, the stringer rivet holes may not align perfectly
with the skin holes because of a permanent "set" in
the stringer. If this is apparent, replacement of the
stringer will usually restore the original strength
characteristics of the area.
NOTE
Wrinkles occurring in the skin of the main
landing gear bulkhead areas should not be
18-4

Change 2

considered negligible. The skin panel should
be opened sufficiently to permit a thorough
examination of the lower portion of the landing gear bulkhead and its tie-in structure.
Wrinkles occurring on open areas which disappear
when the rivets at the edge of the sheet are removed,
or a wrinkle which is hand removable, may often be
repaired by the addition of a 1/2 x 1/2 x .060 inch
2024-T4 extruded angle, riveted over the wrinkle and
extended to within 1/16 to 1/8 inch of the nearest
structural members. Rivet pattern should be identical to the existing manufactured seam at the edge of
the sheet.
18-55. REPAIRABLE DAMAGE. Fuselage skin repairs may be accomplished in the same manner as
wing skin repairs outlined in paragraph 18-18.
Stringers, formed skin flanges, bulkhead channels,
and similar parts may be repaired as shown in figure 18-4.
18-56. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Fuselage skin major repairs may be
accomplished in the same manner as wing skin repairs outlined in paragraph 18-19. Damaged fittings
should be replaced. Seat rails serve as structural
parts of the fuselage and should be replaced if damaged.
18-57.

BULKHEADS.

18-58. LANDING GEAR BULKHEADS. Since these
bulkheads are highly stressed members irregularly
formed to provide clearance for control lines, actuators, fuel lines, etc., patch type repairs will be
for the most part, impractical. Minor damage consisting of small nicks or scratches may be repaired
by dressing out the damaged area, or by replacement of rivets. Any other such damage should be
repaired by replacing the landing gear support assembly as an aligned unit.
18-59. REPAIR AFTER HARD LANDING. Buckled
skin or floorboards and loose or sheared rivets in
the area of the main gear support will give evidence
of damage to the structure from an extremely hard
landing. When such evidence is present, the entire
support structure should be carefully examined and
all support forgings should be checked for cracks,
using a dye penetrant and proper magnification.
Bulkheads in the area of possible damage should be
checked for alignment and a straightedge should be
used to determine deformation of the bulkhead webs.
Damaged support structure, buckled floorboards and
skins, and damaged or questionable forgings should
be replaced. Landing gear components should be replaced and rigged properly.
18-60. REPLACEMENT OF HI-SEAR RIVETS.
Hi-shear rivet replacement with close tolerance bolts
or other commercial fasteners of equivalent strength
properties is permissible. Holes must not be elongated, and the Hi shear substitute must be a smooth
push fit. Field replacement of main landing gear
forgings on bulkheads may be accomplished by using:

a. NAS464P* Bolt, MS21042-* Nut and AN960-*
washer in place of Hi-Shear Rivets for forgings with
machined flat surface around attachment holes.
b. NAS464P* Bolt, ESNA 2935* Mating Base Ring,
ESNA LH 2935' Nut for forgings (with draft angle of
up to a maximum of 8°) without machined flat surface
around attachment holes.
*Dash numbers to be determined according to the size
of the holes and the grip lengths required. The bolts
grip length should be chosen so that no threads remain in the bearing area.
18-61. NOSE GEAR WHEEL WELL AND FIREWALL.
The nose gear wheel well is made of stainless steel,
as is the firewall bulkhead. Refer to paragraph 18-17
for negligible damage, and paragraph 18-18 for repairable damage. Stainless steel patches should be
used in nose wheel well and firewall repairs. Any
repairs in these areas will require resealing with
700P. or equivalent compound.
18-62.

BAFFLES.

18-63. CONSIDERATIONS. Baffles ordinarily should
be replaced if damaged or cracked. However, small
plate reinforcements riveted to the baffle will often
prove satisfactory both to the strength and cylinder
cooling requirements of the unit.
18-64.

ENGINE COWLING.

18-65. REPAIR OF COWLING SKINS. If extensively
damaged, complete sections of cowling should be replaced. Standard flush-type skin patches, however,
may be used if repair parts are formed to fit. Small
cracks may be stop-drilled and dents straightened,
if they are reinforced on the inner side with a doubler
of the same material Bonded cowling may be repaired by the same methods used for riveted structure. Rivets are a satisfactory substitute for bonded

seams on these assemblies. The strength of the
bonded seams in cowling may be replaced by a single
3/32. 2117-AD rivet per running inch of bond seam.
The standard repair procedures outlined in AC43 13-1
are also applicable to cowling
18-66. REPAIR OF REINFORCEMENT ANGLES.
Cowl reinforcement angles, if damaged, should be
replaced. Due to their small size they are easier
to replace than to repair.
18-67.

REPAIR OF ABS COMPONENTS.

18-68. Rezolin Kit Number 404 may be obtained from
the Cessna Service Parts Center for repair of ABS
components.
18-69. REPAIR OF GLASS FIBER CONSTRUCTED
COMPONENTS.
18-70. Glass fiber constructed components on the
as stipulated in instructions
aircraft may be repaired
furnished in SK18 2 - 1 2 . Observe the resin manufacturer's recommendations concerning mixing and application of the resin. Epoxy resins are preferable
for making repairs, since epoxy compounds are
usually more stable and predictable than polyester
and, in addition, give better adhesion.
18-71.

BONDED DOORS.

18-72. REPAIRABLE DAMAGE. Bonded doors
may be repaired by the same methods used for
riveted structure. Rivets are a satisfactory substitute for bonded seams on these assemblies The
strength of the bonded seams in doors may be replaced by a single 3/32, 2117-AD rivet per running
inch of bond seam. The standard repair procedures
outlined in AC43. 13-1 are also applicable to bonded
doors.

Change 2

18-5

WING

12 INCH WIDE HEAVY CANVAS

1 X 12 X 30-3/4

30-3/4

X 12 X 48

1 X 12 XII
X 12 X 8

3/4

2 X 4 X 20

L

1-' -1-1/2

5 INCH COTTON WEBBING

1X4
42

/Ilf

^^^^^ N34

NOTEARE IN INCHES
ALL DIMENSIONS
Figure 18-1.
18-6

^.

3/8 INCH DIAMETER BOLTS

30
Wing and Fuselage Support Stands

GRIND

A Or B

MODEL

A

B

C

WING STATION

THRU
U32i6100

2.00
2.00
.79

1.00
1 00
1.00

29.50
29.50
20.00

39.00
100.00
207 00

uIEGINNING
WIT iHk
U20601701

2.00
2k00
.66

1.002
1.00
1. 00

29.50
29 50
20.00

39.100
100.0
207.00

ALL WING TWIST OCCURS BETWEEN STA. 100.00 AND STA. 207.00.
(Refer to paragraph 18-10 for angle of incidence).
MEASURING WING TWIST
If damae has occurred to a wing. it is advisable to check the twist. The following method can be used with
a minimum of equipment. which includes a straightedge (32" minimum length of angle, or equivalent). three
modified bolts for a specific wing, and a protractor head with level.
1. Check chart for applicable dimension for bolt length (A or B).
2. Grind bolt to a rounded point as illustrated. checking length periodically.
3. Tape two bolts to straightedge according to dimension C.
4. Locate inboard wing station to be checked and make a pencil mark approximately one-half inch
aft of the lateral row of rivets in the wing leading edge spar flange.
5. Holding straightedge parallel to wing station (staying as clear as possible from "cans"). place
longer bolt on pencil mark and set protractor head against lower edge of straightedge.
6. Set bubble in level to center and lock protractor to hold this reading.
7. Omitting step 6. repeat procedure for each wing station, using dimensions specified in chart. Check
to see that protractor bubble is still centered.
8. Proper twist is present in wing if protractor readings are the same (parallel).
may be lowered from wing . 10 inch maximum to attain parallelism.

Figure 18-2.

Forward or aft bolt

Checking Wing Twist
Change 3

18-7

MS20470AD4 RIVETS
24 REQD

PATCHES AND DOUBLERS 2024-T3 ALCLAD

6.50 DIA.

PATCH

4.DIA
00

7.50 DIA.

EXISTING
SKIN

DOUBLER

SECTION THRU PATCH
3.00 DIA. HOLE

PATCH REPAIR FOR 3 INCH DIAMETER HOLE
MS20470AD4 RIVETS
16 REQD

22 1/28*
4. 00 DIA.

PATCH

3. 00 DIA.

5.00 DIA.

EXISTING
SKIN
0O

2.

DIA.

HOLE

SECTION THRU PATCH

PATCH REPAIR FOR 2 INCH DIAMETER HOLE
MS20470AD4 RIVETS
8 REQD
8REQD

2. 50 DIA.
(NODOUBLER

EXISTING
EXISTING
SKIN

PATCH
(NO DOUBLER
REQD)
1. 75

1.00 DIA. HOLE

DIA.
SECTION THRU PATCH

PATCH REPAIR FOR 1 INCH DIAMETER HOLE
|I ORIGINAL PARTS

i':l

REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

OVERLAPPING
CIRCULAR PATCH
CIRCULAR PATCH

0
Figure 18-3. Skin Repair (Sheet 1 of 6)
18-8

1/4 B

B

1/2 B
SECTION THRU ASSEMBLED PATCH

A-A
EDGE MARGIN = 2 X RIVET DIA.
PATCH -

2024-T3 ALCLAD

DAMAGED AREA
EDGE MARGIN = 2 X RIVET
DIAMETER

A^^^
' /-

-- ^

D<
.^\

1/2" RADIUS -- . .;'ii ji.i....iii..i- i/2." R..p.

\

RIVET SPACING =

^'

6 X RIVET DIA.

iii:
- iil:L`.i
.
EDGE MARGIN = 2 X RIVET DIA.

DOUBLER -

-ORIGNL

PATS

ORIGINAL PARTS

LJ^

2024-T3 ALCLAD

IOVERLAPPING REC-

ITANGULAR PATCH

I

REPAIR PARTS
......
REPAIR PARTS IN CROSS SECTION

Figure 18-3.

RIVET

TABLE

SKIN GAGE

RIVET DIA.

.020
.025
032
.040
.051

1/8
1,/8
1/8
1/8
5/32

Skln Repair (Sheet 2 of 6)
18-9
18-9

B

1/4 B

PATCH

EXISTING SKIN

DOUBLER

1/2 B

NOTE
For optimum appearance and
airflow, use flush rivets, dimpled skin and patch, and countersunk doubler.

SECTION THRU ASSEMBLED PATCH

A-A
EDGE MARGIN = 2 X RIVET DIA.
PATCH -

2024-T3 ALCLAD

1/2" RADIUS

EDGE MARGIN =
2 X RIVET DIA.
CLEAN OUT
DAMAGED AREA

RIVET SPACING =
6 X RIVET DIA.

EDGE MARGIN =
2 X RIVET DIA.

DOUBLER ALCLAD

2024-T3

-

1/2"

RADIUS

1/2 RADIUS
ORIGINAL PARTS

FLUSH RECTANGULAR PATCH
(CIRCULAR FLUSH PATCH IS
SIMILAR)

REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-3.
18-10

Skin Repair (Sheet 3 of 6)

RIVET TABLE
SKIN GAGE
025
032
.040
.051

RIVET DIA.
1/8
1/8
1/8
5/32

NOTE
DOUBLER
mDO~~UBLER~

EXISTING
SKIN -

/

A

PATCH

, @21

Countersink doublers, and
dimple skin and patch.

*
<

A-A

/

e^
A^ l

sc



S

# 40 (.098) HOLE
REQD)

^'(10
's~

Z ,_
r,

Parts are available from the
Cessna Service Parts Center.

\XWING

[

SKIN (REF)

S-225-4F COVER

Xs^^^^

<^^^

_~

S-1022Z-8-6 SCREWS

MS20426AD3 RIVETS
PRECAUTIONS:
1. Add the minimum number of access holes necessary.
2.
Any circular or rectangular access hole which is used with approved optional equipment installations may be added in lieu of the access hole illustrated.
3.
Use landing light installations instead of adding access holes where possible. Do not add access
holes at outboard end of wing; remove wing tip instead.
4.
Do not add an access hole in the same bay where one is already located.
5.
Locate new access holes near the center of a bay (spanwise).
6.
Locate new access holes forward of the front spar as close to the front spar as practicable.
Locate new access holes aft of the front spar between the first and second stringers aft of the
7.
spar. When installing the doubler, rotate it so the two straight edges are closest to the stringers.
8. Alternate bays, with new access holes staggered forward and aft of the front spar, are preferable.
A maximum of five new access holes in each wing is permissible; if more are required, contact
9.
the Cessna Service Department.
10. When a complete leading edge skin is being replaced, the wing should be supported in such a
manner that wing alignment is maintained.
a.

Establish exct location for inspection cover and inscribe centerlines.

b.

Determine position of doubler on wing skin and center over centerlines.
locations and drill to size shown.

c.

Cut out access hole using dimension shown.

d.

Flex doubler and insert through access hole, and rivet in place.

e.

Position cover and secure using screws as shown.

Figure 18-7.
18-24

Access Hole Installaton

Mark the ten rivet hole

NOTES:
1.

Dimple leading edge skin and filler material; countersink the doubler.

2.

Use MS20426AD4 rivets to install doubler.

3.

Use MS20426AD4 rivets to install filler, except where bucking is impossible.
Cherry (blind) rivets where regular rivets cannot be bucked.

4.

Contour must be maintained; after repair has been completed, use epoxy filler as necessary
and sand smooth before painting.

5.

Vertical size is limited by ability to install doubler clear of front spar.

6.

Lateral size is limited to seven inches across trimmed out area.

7.

Number of repairs is limited to one in each bay.

Use CR162-4

1" MAXIMUM RIVET
SPACING (TYPICAL)

DOUBLER NEED NOT
BE CUT OUT IF ALL
RIVETS ARE ACCESSIBLE
FOR BUCKING

MINMUM EDGE
MARGIN (TYPICAL)

TRIM OUT DAMAGED AREA

REPAIR DOUBLER
2024-T3 ALCLAD

!-!"_

a

C

ORIGINAL PARTS
REPAIR PARTS

Figure 18-8.

/'

.040" THICKNESS

THICKNESS
--- SAME
FILLER
MATERIALLEADING EDGE SKIN
2024-T3
ALCLADSAME
THICKNESS
AS
SKIN

Leading Edge Repair Applicable to Aileron, Flap, and Wing
18-25

GENERAL NOTES
1.

Balance control surfaces in a draft-free area.

2.

Place hinge bolts through control surface hinges, and position on knife edge balancing mandrels.

3.

Make sure all control surfaces are in their final flight configuration: painted (if applicable),
trim tabs installed, all foreign matter removed from inside of control surface, elevator trim
tab push-pull rod installed, and all tips installed.

4.

Place balancing mandrels on a table or other suitable flat surface.

5.

Adjust trailing edge support to fit control surface being balanced while center of balancing
beam is directly over hinge line. Remove balancing beam and balance the beam itself by
adding washers or nuts as required at end opposite the trailing edge support.

6.

When positioning balancing beam on control surface, avoid rivets to provide a smooth surface
for the beam, and keep the beam 90° to the hinge line of the control surface.

7.

Paint is a considerable weight factor. In order to keep balance weight to a minimum, it is
recommended that existing paint be removed before adding paint to a control surface. Increase
in balance weight will also be limited by the amount of space available and clearance with
adjacent parts. Good workmanship and standard repair practices should not result in unreasonable balance weight.

8.

The approximate amount of weight needed may be determined by taping loose weight at the
balance weight area.

9.

Lighten balance weight by drilling off part of weight.

10.

Make balance weight heavier by fusing bar stock solder to weight after removal from control
surface. The ailerons should have balance weight increased by ordering additional weight and
gang channel, listed in applicable Parts Catalogs, and installing next to existing inboard weight
the minimum length necessary for correct balance, except that a length which contains at least
two attaching screws must be used. If necessary, lighten new weight and/or existing weights
for correct balance.

BALANCING BEAM
-Mark

graduations in inches.

Four-fout length of extruded channel

Grind weight to slide along beam, grind
ends to obtain exactly one pound, andabricate
mark center of weight.

\
/

vertically adjustable
trailing edge support that will
slide along beam.

Attach knife edges and
mark at mid-point.-

Figure 18-9.
18-26

Control Surface Balancing (Sheet 1 of 3)

BALANCING MANDREL

18

After locating trailing edge
support, balance by adding
washers and/or nuts.

Place directly over hinge
line of control surface.

RUDDERS AND ELEVATORS
Adjust vertically until beam
parallels control surface
chord line (except ailerons).

BALANCING
MANDREL

90
A balance in this range is "underbalance.
A balance in this range is "overbalance."Refer to chart for correct range of underbalance or overbalance for a specific control surface.
BALANCING MANDREL

Figure 18-9.

-

Set control surface on balancing mandrels,
with hinge bolts resting on mandrels. Position balancing beam with mid-point directly
over, and 90 to, hinge line.

Control Surface Balancing (Sheet 2 of 3)
18-27

AILERONS

DETAIL

A-A
HINGE LINE

HORIZONTAL PLANE

85"

Balance aileron inverted, with trailing edge
at point opposite cut-out for push-pull rod
85" below hinge line horizontal plane.

Figure 18-9.
18-28

Control Surface Balancing (Sheet 3 of 3)

CONTROL SURFACE BALANCE REQUIREMENTS
NOTE
Unpainted values are not limits which must be met. They are given as guides, in order that the unbalance of the control surface in the final aircraft configuration may be predicted. If the control surface in the unpainted condition falls within the unpainted limit, the mechanic may feel confident that
the control surface will be acceptable after painting. However, if the surface in the unpainted condition exceeds the unpainted limit, the balance must be checked again after final painting to assure
that the control surface falls within the painted unbalance limit. Refer to GENERAL NOTES on sheet
1 of figure 18-9 for specific conditions.
DEFINITIONS:
UNDERBALANCE is defined as the condition that exists when the control surface is trailing
edge heavy, and is symbolized by a plus (+).
OVERBALANCE is defined as the condition that exists when the control surface is leading
edge heavy, and is symbolized by a minus (-).
NOTE
The following applies to the landplane/floatolane except as noted.
NOTE
The "Balance Limits" columns list the moment tolerances within which the control surface
must balance. These tolerances must never be exceeded in the final flight configuration.
CONTROL: AILERON

PAINTED (Inch-Pounds)

UNPAINTED (Inch-Pounds)

BALANCE LIMITS

BALANCE LIMITS

0.0 to +3.0

0.0 to +2.3
CONTROL: RUDDER

PAINTED (Inch-Pounds)

UNPAINTED (Inch-Pounds)

BALANCE LIMITS

BALANCE LIMITS

Landplane

-1.87 to +1.50

| Landplane

-2.85 to 0.0

Floatplane

0.0 to + 7.25

Floatplane

0.0 to + 6.0

CONTROL: RIGHT ELEVATOR
PAINTED (Inch-Pounds)

UNPAINTED (Inch-Pounds)

BALANCE LIMITS

BALANCE LIMITS

0.0 to +12.1

0.0 to +8.5

BEGINNING WITH 20602928
0.0 to +5.5
CONTROL:
PAINTED (Inch-Pounds)
BALANCE LIMITS
0.0 to +12.1

LEFT ELEVATOR
UNPAINTED (Inch-Pounds)
BALANCE LIMITS
0.0 to +8.5
BEGINNING WITH U20602928
0.0 to +5.

Figure 18-10.

Control Surface Balance Limits
Change 3

18-29

CUT OUT DAMAGED AREA

^^by

>

.^^>^

?<

^^^
d%-/'AS''^
^^^

\-PATCH MAY OVERLAP

OR B~
E INSERTED UNDER
^KEXISTING AILERON SKIN

A
~~

|ORIGINAL PART
REPAIR PATCH IN CROSS SECTION

v

A-A
Figure 18-11.
18-30

Corrugated Skin Repair

-

NOTES
Use rivet pattern at wing station
23.53 for repair from wing station 23.53 to wing station 85. 62.
Use rivet pattern at wing station
100.00 for lap splice patterns
from wing station 100.00 to 190.
00. Refer to figure 1-2 for wing
stations.
Use rivet spacing similar to tUYe
pattern UaItwing station 100.00
with the number of BB4 dimpled
rivets at leading edge ribs beOF CR2248-4
*o
tween lap splices as shown:
OF CR2248-4
*NO.
4
STATION NO. OF BB RIVETS DIMPLED RIVETS
22
18
118
18
15
136
13
11
154
12
10
172
12
10
190

'.
/

W
?j

NO. OF CR2249-4
RIVETS
27
23
17
15
15
EXISTING
TACK RIVET

PATCH-

~~~~~PATCH"---'J

~- EXJSTING

RIVET PATTERN

TYPICAL LEADING EDGE SECTION

4'

* NOTE

The Bulbed Cherrylock rivets listed may be substituted for BB4 dimpled rivets
in inaccessible areas, provided the number of rivets installed is increased proportionately. Blind rivets should not be installed in the wing spar.

Figure 18-12. Bonded Leading Edge Repair
Change 2

18-31/(18-32 blank)

SECTION 19
EXTERIOR PAINTING

NOTE
This section contains standard factory materials
listing and area of application. For paint number
and color, refer to Aircraft Trim Plate and Parts
Catalog. In all cases determine the type of paint
on the aircraft as some types of paint are not compatible. Materials may be obtained from the Cessna
Service Parts Center.

MATERIAL

NO/TYPE

PAINT

ACRYLIC
LACQUER

PAINT

EPOXY
PAINT

PRIMER

ER-7 WITH
ER-4
ACTIVATOR

AREA OF APPLICATION
Used on exterior airframe.

Used on the nose gear fairing on the P206 thru 1970 models
and the U206 on 1969 models.
Used with acrylic lacquer.

P60G2 WITH
ACTIVATOR
THINNER

T-8402A

THINNER

T-3871

Used with epoxy (Du Pont).

THINNER

T-6487

Used with epoxy (Enmar).

SOLVENT

#2SOLVENT

Used to thin acrylic lacquer and for burndown.

Used to clean aircraft exterior prior to priming.

NOTE
Do not paint Pitot Tube, Gas Caps or Antenna Covers
which were not painted at the factory.
NOTE
When stripping aircraft of paint, use caution
to avoid stripper coming in contact with ABS
parts.

Change 3

19-1

19-1. INTERIOR PARTS (Finish Coat of Lacquer)
a. Painting of Spare Parts.
1. Insure a clean surface by wiping with Naphtha
to remove surface contamination.

CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
2. After the part is thoroughly dry it is ready
for the lacquer topcoat. Paint must be thinned with
lacquer thinner and applied as a wet coat to insure
adhesion.
b. Touch Up of Previously Painted Parts.
1. Light sanding is acceptable to remove
scratches and repair the surface but care must be
exercised to maintain the surface texture or grain.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.

adhesion.
b. Touch Up of Previously Painted Parts.
1. Lightly scuff sand to remove scratches and
improve adhesion.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.
CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
3. Apply a compatible primer - surfacer and
sealer.
4. After the part is thoroughly dry it is ready
for the topcoat. Paint must be thinned and applied
as a wet coat to insure adhesion.
NOTE
Acrylic topcoats can be successfully spotted in.

CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
3. After the part is thoroughly dry it is ready
for the lacquer topcoat. Paint must be thinned with
lacquer thinner and applied as a wet coat to insure
adhesion.

NOTE

NOTE

19-3. EXTERIOR PARTS (Epoxy or Polyurethane
Topcoat)
a. Painting of Spare Parts and Touch Up of Painted
Parts.
1. Lightly scuff sand to remove scratches and
improve adhesion.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.

CAUTION

Lacquer paints can be successfully spotted in.

Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.

19-2. EXTERIOR PARTS (Acrylic Topcoat)
a. Painting of Spare Parts.
1. Lightly scuff sand to remove scratches and
improve adhesion.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.

3. Apply a primer compatible with Epoxy or
Polyurethane topcoat.
4. After the part is thoroughly dry it is ready
for the topcoat.
NOTE

CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
3. After the part is thoroughly dry it is ready
for the topcoat. Paint must be thinned with appropriate acrylic thinner and applied as a wet coat to insure

19-2

Change 3

Epoxy or Polyurethane topcoats cannot be
successfully spotted in - finish should be
applied in areas with natural breaks such
as skin laps or stripe lines.
When painting interior and exterior polycarbonate
parts, or where the part material is questionable, a
"barrier primer" should be applied prior to the Enamel, Lacquer, Epoxy or Polyurethane topcoat.

SECTION 20
WIRING DIAGRAMS
12 - VOLT
TABLE OF CONTENTS
Page
Navigation Lights ...........
20-33
Flashing Beacon Light .........
20-34
D.C. POWER
Flashing Beacon Light (Floatplane) .
.. 20-35
Battery and External Power Systems ..
. 20-2
Flashing Beacon Light (Floatplane) . ..
20-36
Battery and External Power Systems . . . 20-3
Electroluminescent Panel .
......
20-37
Split Bus Bar ..........
. 20-4
Post Lighting
..
.........
20-38
Alternator System ..
..........
20-5
Post Lighting .
............
20-39
Split Bus Bar. .
.
.............
20-6
Control Wheel Map Light .
.....
20-40
Alternator System ...........
20-7
Control Wheel Map Light ......
. 20-41
Alternator System .
.....
.
20-8
Control Wheel Map Light ..
.....
20-42
Alternator System ...........
20-9
Control Wheel Map Light .
......
20-43
IGNITION
Control Wheel Map Light ......
20-44
Ignition System.
...........
20-10
Skydiving Signal Light
.
...
. ..
20-44A
ENGINE CONTROL
Landing Lights ....
....
. .20-45
Starter
..
...........
20-11
Landing and Taxi Lights.
........
20-46
Starter .
..............
20-12
Landing and Taxi Lights.
.......
20-47
FUEL AND OIL
Flashing Beacon Light .
. .
20-48
Fuel Pump System .
......
. 20-12A
Flashing Beacon Light ......
.
20-49
Fuel Pump System .....
....
20-13
Electroluminescent Panel
......
.. 20-50
Fuel Pump System .
........
.20-14
Electroluminescent Panel ........
20-51
Oil Dilution System .
..........
20-15
Instrument Lights ...
..
...
20-52
Oil Dilution System ..........
.20-16
Instrument Lights ..........
.20-53
ENGINE INSTRUMENTS
Instrument Lights .
....
......
20-54
Cylinder Head Temperature
.....
.20-17
Post Lighting ..........
.
20-55
Fuel Quantity Indicator .
........
20-18
Post Lighting .........
.
. 20-56
Hourmeter.
..............
20-19
Post Lighting .
........
.20-57
Fuel Quantity Indicator .
........
20-20
Wing Tip Strobe Lights .....
. .20-58
FLIGHT INSTRUMENTS
Wing Tip Strobe Lights .
.....
.20-59
Turn and Bank Indicator and
Wing Tip Strobe Lights .
.....
20-60
Gyro Horizon Indicator .......
20-21
HEATING VENTILATION AND DE-ICE
Stall Warning System (Non-Heated)
. .. 20-22
Cigar Lighter ............
20-61
Brittain Wing Leveler
.
...
20-23
Heated Pitot and Stall Warning ......
20-62
Turn Coordinator .......
. .20-24
Heated Pitot and Stall Warning ......
20-63
Turn and Bank Indicator .......
.20-25
Heated Pitot and Stall Warning ......
20-64
Stall Warning System (Non-Heated) . . .20-26
CONTROL SURFACES
Encoding Altimeter ..........
20-26A
Wing Flaps ..............
20-65
MISCELLANEOUS INSTRUMENTS
Wing Flaps ......
..
.
.20-66
Clock .
.........
. ...
20-27
Wing Flaps
.....
....
..
20-67
LIGHTING
Electric Elevator Trim .........
20-68
Dome and Courtesy Lights. ......
.20-28
Electric Elevator Trim ...
..
.
20-69
Dome and Courtesy Lights. ......
.20-29
Wing Flaps ........
.
..
.. 20-70
Instrument Lights ...........
. 20-30
Wing Flaps ..
...........
20-70A
Landing Lights.
......
......20-31
Electric Elevator Trim ......
.
20-70B
Navigation Lights .
...........
20-32
Electric Elevator Trim .....
.
.20-71
24 - VOLT
D.C. POWER
Battery and External Power System
Split Bus Bar ...........
Alternator System (60 AMP) .......
Alternator System (60 AMP) .......
IGNITION
Ignition System
............
ENGINE CONTROL
Starter System ............
FUEL AND OIL
Fuel Pump System ..........
Fuel Pump System ..........
Oil Dilution System ..........
ENGINE INSTRUMENTS
Cylinder Head Temperature ......
Fuel Quantity Indicator ........
Hourmeter
.............
FLIGHT INSTRUMENTS
Turn Coordinator ............
Turn and Bank ............
MISCELLANEOUS INSTRUMENTS

.

.20-72
.. 20-73
20-74
20-75
.

.20-76
20-77
.20-78
.20-79
.20-80
.20-81
.20-82
.20-83
20-84
.20-85

Encoding Altimeter ...
....
.
20-86
Clock .
.........
.
. 20-86A
Ammeter ..........
20-87
LIGHTING
Dome and Courtesy .
......
. 20-88
Dome and Courtesy ...........
20-89
Navigation Lights
..
.......
.. 20-90
Navigation Lights ...........
20-91
Flashing
Beacon Light (Floatplane)
..
20-92
Flashing Beacon Light (Floatplane)
...
20-93
Control Wheel Map Light ..
...
. 20-94
Control Wheel Map Light .......
20-94A
Skydiving Signal Light. ........
.20-95
Landing Lights .
..........
20-96
Landing and Taxi Lights .......
20-97
Landing and Taxi Lights. .......
.20-98
Flashing Beacon Light ......
.20-99
Electroluminescent Panel .......
20-100
Electroluminescent Panel .......
20-101
Instrument Lights ........
. 20-102
Instrument Lights
.
......
.
20-103
Change 3

20-1

TABLE OF CONTENTS (Cont.)
Instrument Lights ..
...
.
Post Lighting. ..
. .........
Post Lighting. .
. . . . . . . .
..
Post Lighting. ...
... ...
Wing Tin Strobe Light.
.
Wing Tip Strobe Light. . .
.
HEATING. VENTILATION AND DE-ICE
Cigar Lighter
.......
.

24 - VOLT
20-104
20-105
. 20-106
. 20-107
20-108
2.
20-109
20-110

Heated Pitot Tube and Stall Warning ..
20-111
Heated Pitot Tube and Stall Warning . . 20-112
Wing Flaps. . . . . . . . . . . . . . 20-113
Wing Flaps. ........
....
20-114
Electric Elevator Trim ....
....
. 20-115
Wing Flaps .
. .........
20-116
20-116A
........
.
Wing Flaps
Electric Elevator Trim .......
. 20-117
Electric Elevator Trim ........
20-1181

Mx20-4

40

°
:

A

0.4

I

-

~ ~~~

~~~~I

Qv

J

~

*

~

~

0~~~~

;_

^r":^^

0~~~~

u~~~u..a-iH
b~~-~ . tJ^'^
^A-:

0^~~>.-,

~

^

n
rf'
«-~~~~~~~~~~

1

~~~~9
4--fi^

3

?

2

0'a
I0_

oec 2 1

0

0~~~~~~~~~~~~

-

^

rr

0

"

20-4~~~~~~

.

~

~

~t

~

i

j
{

~~~.

,

R
!<

m

O
§

h

|l 4£18

.-

1

j

'
_

*

a;

. 3
-

i ir
c
0-w

-

-^

^

A,~~

-iqI?

'
/ '
"1
"';

' .
.1~
-

1

-

® O~~~~~~~~~~,

6

r
:

fe

^^ill^
U^

~-~"

--

-^^

T

'^/-

~.

0.is

.

I*
.-

II

iiir!

i -

2
i

1i tai=
It

i, 7u
14Uo

V

-2.

Z·h

l

I^^4- aiii
Al

,0

5.
" ®^a~l
? ti 1sZ
^--jI'
^ ls
r

^a^

m

CF-

-r

20-6

0

*

!

00SS.^

'

^n~

s0~

Change 3

-%1

~i?

in\^

----

Od~tt,

44

I

.

?? i ~

o~g
( 4 on 4

[

0

I0L.

-Vi

1!3§

.

Itl~
r T

^SSI
I~iIII

e

:1

,-...I-

^trt

s:r.
=5

:,-t'

U UO

5i

c

0

;

r

0-,-

0.m4

i ,e
.' ..

!

U

1.

gS^
Q

<

0

>
1a1~6~/ra

._

I_
1

'a0

*

0

_),:_
~
,.

13X

u"cjCo
<

"-<

S

-CX~

i

I/

--

I

r~

A

1s

@r

l

x

'

^

do

's -l

n

'

2

C

I1
OD-

-

9-

r,

S

4

U)
il

*1
6

,-0C 0

^g oIr

'o

2-

s;~~~~~~

0

P

EdJ^

co

Co,
ssei~ssh
.:

"~ C

!

rn

Lll 011
0 a2

j

232~~~~~~~~~~~~~~~~~~~~~~~~-

~~~~~~~iii

xi

~t~~~~

;

Change 1

,

L'cC

S0-

Mg

'

iS

o9

8
a

20-7

t

a

20-8

Change !

20-8Chng 1

r^^

T ' 1^
-r

I '- Z5
)s-

® eS
~~~i

a'

a

l

~i

'II~~~I

5§t

^-~

I t-r_-"
i^

siS

k1~

5s

8

i

I

, ,o n

I

LON

"

a

; tl (l
^;

Q

iiih

i S.'
s3 ||

1

"A
SIt

W~,

of

,m
hi,

.

;'

La .-M3

..

*; r. 5 J'

?M

^

„

,-

~

o.oj'~

;m*,',

m

0Ln

„

a

~,

?; S S (G ?, .^ -3

S

8

^

I

5

zoo

n rv

1 ,,,..,

1

o

2

"

,

H~~~~~~Change
0
-U
hi
C
H~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~a
" S
So~~~~~~~q.;i!.^ 8^Chage
2

rrtc'

&

0

'

]

,

0-9S

2

^ P

'rPl
UCC 001^

IO 7 /.jl

7
-f

-0 ,Q
4,

: ~
~ N~1~,1 ~1:~'~1~
/LI
.
"

i n . ..

:/ <1
U
ft rLO

4f

--

IIIF
nj-

Ii
2

u

,~01
*
l

r

u
E
lJ
- 1'

_ /

' "7.
£

~-~

,

Ut*

t

'

201

S si t

I

I-o

I,'
I

^

3

U,

!

* *-/
IJJI
~

<

j

inQ
-;in
n

Ii

,

I

dl

l"

2

or

-3

-)

-

| 7

Change 3

9

t

?-

-

L

20-11

£

A

iil

20-sr123

4.37

I

^Ltj 4^

I
0f

Cd

Ch

ta

a4n

0
4

.I
a

'" . *w

zCJI

0

4,
1

a^ :

-

~

h

~l.:
<,
'

/ 1d*
[[. I I:
"

~p

,,,.

0®

ii

ili
0

,

0

^
_~:_~
a-i-~~

s

I' 3.

'II-,!i~s42I ~

'v '*"*"'"I

S£/-3

1tj,

DC

~

J

^r

t

D

~

I~~Ij ;

LI~.

0v

rrh
1

U

i

^.X
2«

~

I'

0

j
j

l

-*

SA

,.r;
I FXio

-fg

"

~~~~
- y
~~6'·~~B

t

g

N

-

a

l/

..

h

Jr<*;

ii

a

I

3

3r

0

=

o

P

-

i

-lll/ '

-\i

Ole

20

I

i

i

I
v

08

-'3"

ll
,

I

blank)

U

NJ -

,n

2 bJa
r/5~

2/2-1Bba

1|

1^

21A2

o20-12A/(20-12B

o

4

I: o
r'-0o0

Nf

a

-

Change 3

o0. .

a

j&<^
i

f j

P: 8-^
i
a~~~~~~~~~~~

^~~B

o~~~~~~~~~~~~~~~~Cag

0i
fi

-,

I

;l;.oI

mI
.

0
^^^^l/S
a

i00
,0< m

,~

0

:

~

CD

.
-I
A

1^

rl~lil~

4

5

"_

j
u(

Lr)

2; 0

ID-

l

2 T'
* w

2
II,1~
I

D

, ,

^-0

__n

®3

I9 11

IIdi

S

®
~~~~~~~0i~~~~~~~~~~~~~~~~~

fl·
-- y ^ ^ pi

B]?~~~~~~~~~~~~~~~~~~~~~

0

Change 1

'

0

-

20-13

er -h"AVT.

ZJ

z

-1

~·5'

1J[

4

A

0

a~r-y~pj

Change 3

08g

gfo HIJ
dw
«. .$f

20-14

W

4

a

a

N

"5

O

Tj,

C

.ai<.

C4 4N

t

-

to0ri

a

Nv(\RR

c

2

!Ld
A

(L~~

L.

:a

0

it.'

"!

.,13i
-n2

3, mJ '-

''5

,UJ.

,

*."

*'i'

LA

.9~=

t

°!,

BP

Chai

Z.

'3i

' I

I4

20-IS

'

^
i^

^p^

2~~~~~0

^s^^

'"

0

e("~l?

C;

N

i

III.. I

a

I I I I

r·,"~~~~~~
COO

J~1 Pis
,

*~~~~~~~~~

~~~~~~~~~~~~~~~

g~~~~~~~~~~~~~~~~~~~

Cn
1 J~~~~~~~~~~~~~~~*L
aoo a
Y £

^'^1

^
1^!^-'

*o

"'«»C

.so

.,a'

"

-.-

0-.-i

>

,,z

iO

*

0'ji

zlJ6O!a
922i-ir

«

.·

;tc

-

I0

.- ^S
3 05s~s°
tuiStdi
z i0i
uaO
,

y

'rs
SAAA
4:"^

i

'.-

o;AA

u II"-

ui1
F:A

W

jy

ooZl]

0

L

-§

-tv-e0

:

a^^|0d'

-

haaeJ32
*...o
A]

0z

^j I i
(

c^§
P^-.'^.a

S

:,,5
7

ZO

q

si.

0^n.|1tQO

-)

,.'--'o
t r~sna
-~

~1

~

20-18
I
Change

o.~

20-18
Chng
~

~

-.

o~q

--

--

oZ

0

i

*0<-.0

II

53 z
0

a

W

:>

o

~~~
0
/\ I

,r^C

I

Irk

, ,

^^-

z a *'t

*u

0„o^is

~0

,

I

Q

oo

ne

!

~

^'/'^/ ' rf

---

~ ~ ~ ~ ~~L

12l:a

-J

O'U'Z
uC

/

b-ffao~~~~~~~~~~~~~~~~~~~~~~~~~~crtf

'°'

20U-1

21

,-,

^-^i1

*

20-20

~

-w:-C

.UJr
<

3
Change
W0-i0
e rha

1

_!®

:

I.~~~~~~~~~~

I

I

.W

>

;

~,

I
N'

i X.F<

02
,,,
_
-!}

r

ys~~j^^ ui ja j

.l

ii
1-'

a

- 3.

T /I"

.

I

^^1

/

.~

'

^CI>

iKt<

8*,. Io

C)

(/-"-

Sd.i

is%

o.
t

Fo 2

I

0

r

ddf

i^

(

------

-

~l:P

I

00

i

-.

.

I I I1I I~1~1
1

*

Mi1f

iit

iif

T

2

1

n

H?^3

N

Chwulc 1

-

-

NU

i

20_21

I-l

'

C

20-21

UE

an

T.3

ct

-

of
f

o „ J- 4

C

o

"tft

ran

"7
O^

-.
ss~~~~~~~-

'Jul0~~~~~~~~~~Chnge

J

"l

*§

a

0

§

4

S S s ? *2S
I

0-3

0.r,

:<

0Ci
1,

I

"ig~jl-0

co

n ^ L'^

^.,

'^

,

^
4

/

Cl1

II,

-

I,.

I I

i,,i1Ii

ml
p

-

^^

· f I'0
ld

5

^

g^

oD

3

0

i'i 4

0

t

(AVi/g;I
---

-rr

§2

o"".I?

aiS&

i$s

~~~~~~~~~~~~~~~I

^'*'"J'^
^

-

z

a of
o

nip)

j.i4.

br

1

'

1

K'^-m~~~~~~i,"

511

0

M

,f $
ii,^1

¶
\

) i~~
n

'~~~

r€
2

;

8199~~~~~~~~~~~~~~~~~~~~~~~~~~~0

^-ff~~~~~~~~~~~~~~hne

l'
A

'P

uJJ

,.-LfcJS~~~~~~~~~
-;i

a

Zzd
'-'AAY
Sm

-

2

I- ,

L

o
o^

0

-4

8

jj

0

)

1

-L

0

4

Ss>~~~~I

A

S~~~~~~~~~~~~~~~~~~~~~~~~~~-4

- [4^

^ |

1

&
~~~~~~~~~~~~

zl

3~~~~~~~~~~~~~~~~>02

0." S

w(

a-^Otl~~~~~~~.4

J(

""M

|g
rj

02

It
00

s k,-;

00
0 o

'^-^

~

01iO

/

20kP
2,

20-26~~~*Change* 21

VT

L

0~~~~~~~~~~~~~~~~~~~~~~~~~~

J

~

~

~

~

~

I

I

~

II

~

~~~I
^^

§.
.

A
-- fr

0
a~i

jl
^

0-0

' .

CJ
Q

)

^ °'*

~~~~~~~~~J
:

a~~~~~~~~~~~~~~~~~~~~~~~~1

4''
'^

If)

;
-

r:

a

-

u ti S
mI

o

K N2
11 ^ ^~~~~~~,
u

u-'n
Is''*"^

wm
l

l

As
-

. Is

1 3

-

aza -a

'E S

~~~~

^

-Dts

at
34

-

I~~
il3
I
w 0-g
S
> 'HU S ^ H ^^|/1|dI

-t,4 4~

3.

C)NII
Tta,.tf

;oo
4ri

~ ~ ~~~a1 5

a~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~0

~ ~ ~~ ~~ ~ ~

~

5~~~~~~~~~~~~~~~~~~~
~ /- ~ ~~
~
~
~ ~ ~ ~ ~
u,
~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~

-^

>**u~a 1

~

Q~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~
0 0
0~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~

4;

I-AN,9

Z*Cag

'

a

Lc 1<~ "

U

0 7..

4%

900sj

I

u~

1/

a^

°

4
0

..

a

^'

t

'i
;~@'

®,T r,

6
v'^
\\ .
\"I
r^

0I

,J

ijri^

v

*I

L

-1

y2

f?

ii

»
^

i

.

r

I
I

"

z

g
SnS

C

j2

I
^"-

J

,

±

;

20-27

20-27a

i

III
-o

1

~~~~~~C.e
C
1

I I ifl^^.nirt
Tn

-" ^-*-1 n:
-- c/5 SV -- ^ -_ ,ji,-a

0,

5

-

^

-

„

-" f

n

;

00

I*

20-28

;

ot-

.o

^^a

0-2
uJ

6l·l'

- 3Z

j

4__ -

Change I

^

I

a04

^iD a T

s

5

Li
9?O

Cag

i

L-

II J,

I

r

®

^

°

£

(2

^^~~~~uV~~~~~~~~?

1

,

!

-

jVIo

0

1di
,

2 c-

1

tzl

li- a

o"I

5

/

n

0

,'

e

^oI

A

9

t
D...

ta I

2'

A-

-

0

>

< 0

-

cr

A

.m

. *-

In

0

I
'

'h

,o-

-.

d.i

0o

I1,

-_'

«m

,c~~J
s ,
.~

-ID
P:

i

?Y

^

Lwld~s~a~

0

,o¥

°

00000000£OW
S

1

n o
EF/

I

-- --^

CD -i-(r"
,jr-a
s®®'s
®(
*
~
~ ~~~~~~~~~:"~~~~~~~~~~
~~~~~~~.J~~~~~~~
~~~~L

-

\

w~~~~~~~~~~~~~~~~~~~~~Cag

3

^5

^

a

I

<
(9

C

2-3

,

a

.A ,

IN-

iJ
s-h

-I
I
IMZ=~-

0·

3

2-<

20-36

i

^

I,

4 .ei*F00

4^ ^

uI2I

CI
ChangeD

Change 3

^

*v^

jaar

J_

-- -ii hcj
Lwr &u

Fl

I

mo

6'm " Z t1

'Br~,
~ ~,-i i/iin~ji~i
-, -L /i/'ir'i/'^
/-L^
.4

^j

I

< "I

I3 Change

? if

a

_

i~,*t
a:
~
yiif /'i/'i/ii/'

L

r

-

^

I

¥.1s

§r- |

S

n

P

01

I I2

I-

z~ilj~s~r

.i.

mji
;,t >I^
i 1,1-I
.: 1
It

OI

2

I3

"*^.-i

8

4,

-I,

-·

I

0t

<

^S'^XA^~-

I'

v·v9

-^^'~,^*-~~Ci

a

5,

D

|.

·

ii

-

-_

wt

II

I

20-37

20-37

v¾1 Bl~lnl
s:

C1

£
6'g

IIE ~1,.

xI°0

0

IfJ

'J 01
'

7W
d)L

'°

8<
i»:
o. '

i
.2

I I

i

0

!!
'

,l

kI

"

1

'

0,

s

Il,

>.

tOy
z=zs

0

'

"
c

s05
0

I^OD^ ISLLJ2

6

H~a
ii I
a

i..

z

0

'

~

'v/niraa r e

%{o-

6,00

Z j5ji~~~~~~~~~~~~~~~^

j 0]~ ~~

on~la~h

s

o-. 

i

YO

I0N

I

:::;5

* . '

S 0 '

1- 0 h

F

t jl

ci

a2

a3

I

a

w

r»~~~~~~~~~~~~~~
8
5 Fr- r-w

"'-

?§~

o nn ^ r^ 0 !

. " . -. ^ jl

t 1 <

'

? -

^^

-P^

7

-

-

.*'i

-.

f
i'

'

J.

°s

iS^-*<; o
^

*<

¼'I

-

U

> >- Pv-

n-~~~~~

>
->-^<-^

^^~~14

2

03 <

> t

s

~ ~ ~

a

^~~~~~~~~~~
^

\

j

'C

\

y'l
_

2 ~
~
~
~
~
~
~~~~~~~~[
~
~
ii^^

n~

?
^~~~~~Cag

;

i
i

coa

z

s

"

<

'

0-

'".
- ? ;'

i

_
''0(^]^'

]'J
^
"!'''i?

'~4
2-

-7

-

L~~~

;

(j-

:

(d
<

-

1

^^

^

^

.

v

ia

'

K

i

- U

2

^W

:

'~~~~~~~~~~~~~~~~~

blnk

I-^D3

'-0

.

**-

^ ^*
0:^

\^
!-;

*,'

4-'

204A04f

A

-

00

|f/\_

>

0,^i.

b~~~~~~T\

^r<-^t;i^

XNn54

J

--

i

s

^

Ir

J

2

;

,"

o

0I
J 'o^-'l

J

S

a' lilis.2:~l

!1Il*"llj

sf inrs) ;

*' ^^
^

1i

s

=
n

~i^s
20-45

S^So~~~~~~~~~
ts4

'?o^~~~~~~~~~~~~~~~

^Ji^L-*~~~~~~~~~~~~~~~~~~

§

>/ ^l'~.....I~e
^t+1®.s^^~~~
^ J ;!,,20-^4

^ vf . <^ S^<^^

I\-A
IIIIIJI *^^-"~~Ik
ll"l l 'i I'

~ ~

2

· %SC'4'4.4
dl, '4,\\^

»

a~~~~~f
~~~BN'46~
~
I d2

/--\

i--q u P-^--^r
Jy"~~~~~~~~~~~~Cag

20-46o-re~
Change
~i I

20-46

3

Change 1~s

s

A

A

-4

-1-

?
1,.

t

~

5

j

I

~~~~~~~
r

U.~~~~~~~~~~~~~~t

o~

20-4
St

';

- .
Jr

^ ^^

0

s

S

iI I*

P~~~~~

* "J^
v^

k A

A,

J9^^

Y

~

-

I

ii^ % ?0
ill^

t

^
^.->''

l^^le?

~~~~~~~~~~~~~~~~~~~~~~~~AJf
~~ ~~~

RK<

U'

I?

',

|

'l

~
0

( ~~~~~~~~~~
~O
tO

LO

~~~~~~~~~~~~~~
i

in-

^V'

z

4

8Sl~s

*I~j"J.i

U'i

.a

nI
Oo

a

o -0

1~

8

I

i

"

20-55

;2-

u

Li<"^
u

.et

^--

hi

.

0

114~i

Chn

'-~~~~~~~~~~~~~~~~~~~~~~~~~tnV~
'~"*

Ks

I

^^ s 5|a~~~~~~~~~
ig:,;

I

0'

~~~~~~~~~~~~~~~~~80~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~0
1 1 11
2~11
1 1
1
4,4
~
~
~
~
~
~
~
~
~
~
~
~
?/
^^^
^
^^^^i0
s~^
Q
(^
^^ /^i^
n
~~~~~~~~A

^

wr^

O 'D~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~'
-o
^
*,
,N

-~~~~~~

19w

"Z

^
^~~A
^ -V^N%?^
~~~~~~~~~~~~~~~q
~~~~~~~~~~rd~
~

U1N

u

ztfi^

Trfi/-

Z~I

"^

; ·'¼

: 4a
s

4

^n
t n

o" <0l
um

<

3a
ISP'

Change 1

t

itI

25
S.

....
on?

__

---

0Cha-ng e -

i^5

~2~~

H

s

k

_

I

,?

.1

-

_

I

f

I--41,

1

ya'

S;

eli'

^

9T

v

t

'

%

!^

5

-

t-

CL

3-

-

I~~~~~~~I

.

w

!is

,

(D-

'g^S£D -- ~

r/7Y^

O

. ,Co
_a

Q

--

(UDD

,-5i

D

°°

-

1

-,-----.

.

1 ,

klil

-.

i060-982-01

1

* illtl
g8

.i

i

LJf-

i

S

Pli

u

ii¼

^S ssor

-

O°;

i ii

a~~

.

ggp^ s&; ?? ^r ' /I

05

=

20-57

7o
3 25e

Change 3

o

00,

0

;!

a8a.
'?
-

i~1~

IH Io~-I

i

*/

~o~1~,.~1~
-^ r

20-58I
C
hang
I
0-~O-

Ciiii
:^

£ J

Vs«

t

"Ii,
«f

!z A

-J

i

Z^

z~l^tj

naS
6§5
F*2^
°^> -/-^

°
E 3

1?,o
>Z,

^-

z

ta
* 3i u 11
?:3~~~~~

3

·

S

:

B

i

s gIol I£1~1
<
I~':
I®

^y

5I

...
~

LJi*t

. 7,~
a

y_~~~~

G
I

f

j

Cs~~~~ -^

FI

I

OT

oA ^ "

FU

Vfli~m

"

I

I I I I I I

i---

ASOLO~~~//

'-

I I,

hB
v1

_0I" ®
I=1'11

is

I

L"·L'

I
n. t

ZI
, I
a

z0

"-1V'I
O
N

9

-

I,

?
U
?

l,'
WI
4

w - u~~~~~j
I* I9<^~1111

r30
10
a
L

-

i d

i;s.

I0

,A

&oI 0,
,-~/i
o

0-

o

.-

-

i. I

'

^dimi^^o* I

*?in
n ii!

00
00 9.ri0

1, ~ tn

a

yr ~

T- s

~~~hi
-^ih

,I

A

-a

-"
^

~
a

1

1
7
^ c0]

5

V

~

r
u

?

>LJ

S?3



I

^<4

is6ra 19 Fi·

*^

·· =

!

0.

'

i^|

1

C

I

^

L

,

tf

i '*
;

.rzr

t-

--t

I

/

-

' 0
S. 4I
0 0

- -35^
'i

1*

-l-r-i
5

11

X

"i

^

d3

0

.: 1
^Z

~:<

ooo

Ao;SS^S^^aJ~s~iO £*'^^~

',

Ot o

.

2 Z < 2:

? °: ^ "J."

izJJ'

0

^

^

^^e L
" O^Od)^
~ I ~I ~~~~~~~~I r->J^J'l*,t
^

,

^
u

L

111

-

Sl

O

rZ )

1- -5

O

L

4

o - A S HJ
!

i^,
^

j I

0

6L

^ ao('a
14J f IlhI1 2-I~;~/

ss^

^
a S*SLo

.'

EC

-

Sc

"Si

~oI,,0

,I~r
1
<;

g

20-64
Change
3

I,^^

<

20-84 Change S9

-~~

^

IJq~~~~~~~~

~~

5®<

_

co
20

1

^

03
~ ~

-_

_

*^'

s

/^,^

^

t

~^

~~~~~~~~~~-

'
~~~~~~-64

£y^|il~~~~~~~~

UIJa
I--

C,

(?1?')

-0-rn-cs^'

J / ?/-

i

-~~' O

T

"t0.L5"11u

^^

fAt

O

-

9Or£AA*SA

,;

^

-~

0

?

^S^

2

Ch|6

O

;

Zoo

c D!70

.- z.0

I
I~ ''I >1Z0
0

B

z

BJU

h

1

"i

4 j -i * ^

''

-?<'8

/l

.

ooA~~~~~~oS~ipiftip'"!/'^^

ImI

re

|

,

A

Ss°S~~II

E
» .rtpJ= Q~~~~~~~~~r~~r-J
in~~~~rt ««j- [g~~~~~~l
?B »6~~~
^t-^"~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~I

s!~~~~~~~~~~~~~~~~~~
^ o~~~~~~~~~~~~~~~~~~~~n

(

^Lj

•"

^

'-

-o

:

:

0

1

j 0~

Zj

(?^
a

^

~~

~

£

o

,

sIr

I sm

>

Ci

--

I:m~~~~~~~~~~f

ip60-/I

2

-- @r~

ouk0 <.- H

"

2-0
t <

to

0

20-7

.n

;„

A

-S

U-'-

3

I

%o,
J

CagJ
r,j
^^^3^5^

M^

ha

6kI

s

ay;= L- ^ 5^ .'^
a b

i-

Cs

'n

j

^

c

^7 .

.i:^

^

I"'

9
y ^-^^j^

.,,,

*rrhs-

n

I

A

i:hn 2

AA

n

2

i

z

a

r

t

^

-

c

I,

CD

I
i

za

0.

I

-

- 6/

!

II

-?

s

-n

II
,~

~-

J

I

*~
-o

'Sl~
~t~~
Ir~ll

o*lIg'"'-^~

- -- 1 /" ,~£lJ'r
I£$I'~ B,_

o.

-A

IAI~

_,-S"

n

-~^-^^
-

O1~1-~
.,

-"l-'i -°°°*
o /

-

'1ooooc

^ifOO.M-O%

ll

|
"

11

>

-

Ir-5

-

~
a

O1

"

o

®
I i ino/0

u

3

oj

W

S

,

i

-J

0'^ll
r
t
„ (00~~~~~~~

'9

^J1 "' . 1 -'IJ;

N

0,NI

a / ~11
14

,»

tz

I

00ID
a~i~rfift-

'ii

11~T,11°~11~~~1

al

l

Q>.3So

-ar:3in~fl~j-Ogi(Or

07
Change
Ft 3
LUG

§i

rf)
*(L* S
t,'1 ? ^

I~

0

ni

.Ij

___ _

a

L

0rit3

-

--

I

I

gs

8Q

a

-

I tI I

I

c

,

b

EoS"~%

Change 3

o Wj

30

^ tn "

I
.j

t(13

"u,

S

~~~~~~~~~;Z~~~LI'

r

r
?

t

wer0=

*c .

'a

x3

20~7i

5- 1^

20-2 Change 3

l ^
s

a.

c

;

£

iH,^

-..

it

i

_-

r-l'

~

~

z

0tj~

U4
iljiii,-i
,,

^'i4- *

0-ii"?
?t
§

*^*^'.oi~~~
^
S

S

5fr,>i

"? fi

IC

m
II

E

1n

~ ~ ~~~ ~ ~ ~

Ierr,

-^j-^i-^
J,.

__

/uon .5S

!

-

~
^

fi-0

I"

_^||,IM

cu2i

o ,,,

-

;

>

OO

zo
.(

Izi

I0
2

ifUD It
1Os

i IFI

II

t

P-

I

°1/
K

RD

N

F ~
%'

*T

P
'iLL~~~~~~~~~~~

^ -o 1

T

4

N

0

Change 3

n 2

8

20-73

^L
·

N

-ce
4^

^^"^"

\

^ ^""
f

)

1

-^

:

I

-4
I

LDl

2]r

\~~~~~~~~~~~~~~~~~~

I

!~~~~~~a
f~l

- ^r

^
^^--ff'
I.'
I
( r ©-.(
- -'-------i

*-0

'

'w

S

.

-

00
O

i!~~~~~H
t

y

ON
":

S

I

--- w

5rfi
ad

<

"" Ztg

t

-

-

5". hi. i

-

FPChne
J

oo(

I

r=

·

z

9

"

0t-.,

-

8",

y\

(
io

'3~~~~~~~~~~~~

"'"""I

a0

F

<

20-?$

?
JI

0S??

0 03

CHn.,e

iSw

20-76

Qo

._

*W~~

rrtN

Change

" L

«a

0

ei

fl 'A

°1

I
1,

0
t1il ;
3
'LCI
5 :o
.o

I~ '-

0

ic

-

-;
j

25

LI , O

t;

2(Dm

t'I

d_

-

3

as

^i3
gpi

1 s5
S zli
m

il-o

OQ'';oO»,l*'5

o

S uw

S^lI

-I

-^

Ii'S

Sto

I'1J

av.
45

*A;s

L.J0

'0

-

\

IS"t

I

-7

Iy1~

i
c-0n

aJ

-'o~~~~~~~~~~~Cag

/

H

~

~_

of*n

ou

3^

u

s

4

Y

L>

Ji3(9

I

?1

II

320-

3

-[-y~

-J

0

9

r1^

O

.
8i
2

Bt

EZ

.0a

I

Z

L

Is

Iz

r

111
,9 a
nib
-,

"°_ __

I

0a

Ca-

1O -S^z

\J

......

o08

.°u :

'

i

82i

2-8

ii

,i
5 "fJ

20-4

Cag

JO8'

(
.l~

I

Ir^

I0

.1.

0

I

r

1'

0

n

I I
' ** ** *~III

*:<* S

0

CD

1

10

a

*

0

h.

o

0

_____

~

-

^

r^

°<

0

-i.

p-'af-r'5

-S

S

rJ~~~~~~~~~~~

"

C

I I

--

n

I

I

3

I

-i

S; '

~

I

I > ...

-

Q

a
*

' 0

^^

5 ^ B/°

0

dof

9rT5

A

a,o

.;

j.

hl

C

2
pI,

'

O*

-o

:

0

,

.1
0-87o

c,

0-8
2

°

S0 s

f

I -J

^ ?"-/ I sIl ,

»fl

I~~~~~~~~~~~~~~~?S0
~ ~ ~ ~ ~ ~ ~ ~~~~
L/ ~~~ ~ ~
a~~~~~~~~~~~~~~~a.S^
iv~~~~~~~~~~~~~~~~~3r
~ ~~
~
~Ju ~
~
~
~

00

fr~~~~~~~~~~~~~~~~~~~Kr

1

0~~~~~~~~~~~~~~~~~~~~~~~~~~~to^cs^i
Chan=

S~~~~~~~~~~~~(03^;"

20-88

S'a'

0

'0-8
C

3-

Ao..

0i^ac,

Pi

0e1

Change 1

208

-JJ

0

9

S

:;

ChangeI

1

4

I

i

sOO

0

LI

o-

I I*M

D

S

91

rL

i§

~:

toae

<6
O

,!

,<

I

Un.

-

jrdj
m'^^^1?^^'?!

i

irs*>

7 I

oi

"!

35-

3

±

4;.

s

0

^
1/I
/~1

'»?

--

-

^

l

'~
»j
-5

Y,

"
U

6r~~~~~

a*xp.^Jia~~~~

II-~

w^:~IYJ'u~fta

-r~~i?~~

"'

4*-

I

-j

r

^,--

^] .3

^^-^
"
*"0"

In
%9 ;.0'~
O
'

I

W

~

^

1

g

I·0

I

3 8 p

7 ;co\ja

E

>'~~~~
2-0'

Ul^

Lcd
-AJ
5I'
< J-

iU

28

o

-89

*

1

0' L

3

II

5.C3ne 3

0

-

,s

0d ;:~

fO4

5
5

SM

I

I

-^cr-

*

^ „

.I

S

Y--E(u
:-

^

?^

j^g

3Q

F0

5

^l^
~~~~

~ ~
d ^'U^

it

0

^

~(J
r
^ f l-

2
^

0~~~~~~~~J

I DQ

~

o

Ji

4

«

0

^?

'

~

/

t

So^
^a
Qi/iA
4
^ ^

~

£I

~

Is

--

I, 1

'J
8

I

0

09

0aN00

-

C

t

,b

0

> sT

O

j-as

3

J

j

*-

0i

20 91
"*u~~~~~~~~~~~~~

* ;!!?~~~~~~~~~~~~

11

n

Ilis~g6

z

,
I"i
M o^
-

z
Can
i3

1-0

8

~'-7----

~~YI~~~~~~I

j/|

--

I-

p^^^s.^ i!
o

_s

^^Ut

i

8
(yCha
ge 3 2-9

^1^*3^^' 9 *^t*Si

^^

5jS~ ^~
-*

^oo-fS'-^S^^oQ"

w w fJ

tf^^ niif

(^^-u.?

J'fi

'^-J~~~pro^^iri~~~rfirj-

20-9_ C n

0

011

,

P4.---00
11

IS

6t'tt
as

4'

I

t

C

S)

It-

ii

.

r

~

-'

<

II

n, o

-~

1__
^ "S^?:
a
/w

I-'

^ ;'

C\

o

- , 3

Z4

1^/ -

r:

® I ?1

-

010

'I" °"~av
z

r'~,

-.

YEE

I

s

0

z >O
r

"'s

I,'10,-M,,I,,-l,,l,,H
I
!

5,o.~., ,. l

-

r c\r

J00 j aQ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~

'

g~'"

*

~1

&,]gL
*~I,

-lI~

~1

:iTl v

' It

2

j
U

.ir Ilia

4an J

^

Cge

I

iii

I-

:'i

Q~~~~

I

,'

2:

nr.

-J

~:,I~~~~

^4--

",_

U^
^'
^a^i^^LP^^

^q

3

'
?

in

5

^

;

I^S

.

u
-~~~~

"'

|

|/

--

0

**"S' IP4

, I

::* :*

-

-

1

^~

Change -I3

7"

- ^

-,

.D

-

t~
IS

20 -93

-ti

9!
!)
^^sr~~~
r a^ n"-

T

!

-C

I

-<

^-

. S- ^0

^

---

~I

-

,

-1

^\^

^

c20-94

*'

2

,

~

0

-S

,

e

~~u

-

oO
-.

P-

' 5

Change

I

;,,=

,

-

3

ii
^-~Ilel-ldel^l+H

S~i3?3rftf.<.,-*»

N
c

t

g.

0*-~~~d

S--'^§ I9o I

r
Se:
?

S^ g?
^
7:^-'.d r I~1
n'

c9i

Wr

;

0l

R4^,^^
m,- 2!''h~-.

Ce

iwn

S 2 z"'I

C)l

z~d

S *'

.

t

"·3'

Y

o

t
.1

g

Y^

*

~--Ioi
1-Can
I3'20IrT
0
U

>o
tQ
A®,-O,,®
* !__}oh

®i

1

;

c

.

*

0

oo

.

Q^~~~~~~~~~~~~~~8

%

©I

I©P§S:

n

81
rA^

:

-J

· fII

I

-- T

ow

g

ti -

'/".

0-

10

0

~ 3"

r-

Br

A J

I:I

~

0
Sas.

Q

L

Of,
,,

I

1*1

c

,..
0U

S

3

1i aXv

§
§!

rji'tnM 1-l0A

4~~~~~~~~~~~~~~~~~~~~~ *

©Q^" e^"~~

^~~~~~~~~~~~~~~~~~~~~~~~~~.

-- -»"

I S

~

^ ^-'[- -* - i

I-U¶if
1
O | Ti
'-'
' 'i

NI-

0

f0>f08
IlbO^^J~i
Change

Change
'~0o-108o

2010Caney

S-^5

S~T^

Z

S -

I

raw

?

^^^

C^s-'ll

®T

0V

awi

T o

;IC-

5060

'

z

es?

~~

o~o

01

soo.<

^

0-0

~1S

3

~

~

I

D'AL

k;

i-T

~
i-

i

U
^

'i

a

^O^

~

i

H0

5i

'1

l,

W

IIII

10

I

3

20-109

2

5

n

rLY i^'Lii
2
u

1

C

ii ie i
-

,c

*

< ..'rfl
.
--.
-

/

yl4 ai<

Change 3

-.,

fit

*

0

e4

an&/~~~~~~~~Change

i"~~~~~~~~0

I

i/

^ ~~~~~~~~~~~~~~~~~Cag
~ ~ ~ ~
T1

®i~t~`"

rf

-a-jQ-Oj--t;(yfls;

~lel~; we

i*-

IB"ig5

0

q VI/

z^
''

Do.

"'i

TIa

-

"'V

a

-l'
.

-

-

A

I I

I

,,z,

ih

S

OON

i

20--

.-

dll,

~

^~B
ZF-

hi

I

20-

!

1

3

l

SriSj

Change 3

20~~-<><<1

4,!

-

I'JAi

20-112

f3--I

co

iS

^

ip

A

*S T *>
-

s

-c,25
-,
"-

+

0

NI~~~~~~~~~~~~~~~~~~~~~~~~~I

*I

?F
I~
J 5 ii*~~~~~~~~~~~~~*

*

-'.aa~~Wf
oqLIN,

a

N~~~~~~~~~~~~~~~~~~~
a< ~~
z . i0'i:
I
5r"

d)N

WA

-

y

31i
i~f

-

~~~~~~~~~~~~~~~~jI~
m '
La1II
A3^
.A e ^^,
^ A £ A~J'>-^" --IB
3t

^w
np

>t&.

*Qtt^

amj W

ii

£SS~~~s~~aftaS5$3^ ?3S~~~~~ij»^
.S^S:Qo'ct- ^
^

I

`TT~~~~~~~~~~~~~~~~~~~~~~~>~~~~~

F4
V^;3m~

-

'0

Tl®
'

Ak

rY\ir
%T. os
11c7,

I~~~~~~
I I-i

I ^II

1-^'p
-- «

;1

©-

"

^

u

I

1

fifl<^j
a ~ *»^-ans

- "

Chage
20-114~~~~~~UVU

*j

a?^

Vs
I'*5%j

haz

~a

0f

^

0

?
aA
f

N'

H
P -Ijji

di

I

j

r
j

id11
d
I

20-i5

sR

Sh

e]

jiti

Ch

20-115

aa

Change 1

'p

-z

-J

-

diN

o0I

*

fl

uJr~~~~~~~~~~J

-·0~~,

I~~~~~~~~

L

,

-,

0 -i
ciK

U
-,

¶ii

-

';"1^ntf~[~~a~~l"'"

oT
~ir

7

-- ?
nB-.t z G1

4sJJJ
¢-

L

''7

t

.JJ^J4

-t

n-

3

"

*

aa

-

3

'

*

."'

*oe

.j
- --

t- .$

@1j
0j

0-

l

ac'*".-i^^^j^^^M^'i'^s^bi^^^
azb
-F

3

~Olr
A .*t4

z

r'

20-1 16A/(20-I6B
I
blank)

|

^Chne 2

20-116A/(20-II6B blank)

~

Change 2

1:3?1

'

>~~~~~~a~~~~~~~~~~~~u1
T'

>*5

5 ss
3?
-'-^1-

a--Illl

r:lb'
i

1

9

far
a

.J^

& -W

-

ii,

'5

;:,

>

-

00,

'1

r

f f

QO'C '-

uti,

^.

-

-^'u

W-O

0

o

-

C

-'"

-J-4

A

d

I

*

A

5

°
,l:ll

S

QWIj

v .4 0

o

Z J

Z

'ii

-_,ga
a

i

aa
CJn

I
11§1.

pCse 3

a

te W
j0000
A

Change 3

,

Irll
:
iu.iT

s

S

20N

20-117

4

1-.

201 o

20-118

^

C

I

p
k

-

I; ' I

·-

Jr
"J

.

IE

\ i

ie.F/,

:

5 T;'^,

o

3

I/

I

i

o*
tr

U.

1^

I

/

7

SI.

I

,

i

-,

-

.

LV -J

ri

IT

--

'I--

0

-

4t

I

I

un

""

s a

|d3

U

K)

o

Ii
~-

I II

,'

n Ub S

r^)
A

-.

|)

Ca-

Ti

0

-

-----ft"

-- "-'t I

~

a

--I

i-Y
S

i

s

-z *r

J-

t-

3a3 3 ^n.

2

1r0

?

U

IMI



File Type                       : PDF
File Type Extension             : pdf
MIME Type                       : application/pdf
PDF Version                     : 1.4
Linearized                      : No
Encryption                      : Standard V1.2 (40-bit)
User Access                     : Print, Copy, Fill forms, Extract, Assemble, Print high-res
Creator                         : 
Producer                        : Avantext, Inc.
Modify Date                     : 2007:12:05 15:57:47-05:00
Create Date                     : 2003:03:03 16:28:27-05:00
Title                           : D2007-3-13 - MODELS 206 & T206 SERIES (1969 THRU 1976)
Subject                         : MODELS 206 & T206 SERIES (1969 THRU 1976)
AVTX LPROD                      : CS04
AVTX LLIB                       : MM
Page Count                      : 542
Page Mode                       : UseNone
Page Layout                     : SinglePage
Has XFA                         : No
Mod Date                        : 2007:12:05 15:57:47-05:00
Metadata Date                   : 2007:12:05 15:57:47-05:00
Corruptor                       : http://www.w3.org/1999/02/22-rdf-syntax-ns#li
Author                          : Nobody