D2057 3 13 S 210 & T210 SERIES (1977 THRU 1984) Cessna_210_T210_1977_1984_MM_D2057 Cessna 1977 1984 MM
User Manual: Cessna_210_T210_1977_1984_MM_D2057-3-13
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Cessna
ATextron Company
SERVICE MANUAL
1977
thru
1984
MODEL 210 &
T210 SERIES
Member of GAMA
FAA APPROVAL HAS BEEN OBTAINED ON TECHNICAL DATA IN
THIS PUBLICATION THAT AFFECTS AIRPLANE DESIGN.
REVISION 3 INCORPORATES TEMPORARY REVISIONS 1,2, AND 3,
DATED 1 DECEMBER 1992, 1 APRIL 1993, AND 3 OCTOBER 1994.
COPYRIGHT©1996
CESSNA AIRCRAFT COMPANY
WICHITA. KANSAS. USA
D2057-3-13
(RGI-50-7/02)
10
REVISION 3
SEPTEMBER 1982
1 MARCH 1996
Cessna
A Toxtro CompJny
TEMPORARY REVISION NUMBER 8
DATE 5 April 2004
MANUAL TITLE
Model 210 & T210 Series 1977 Thru 1984 Service Manual
MANUAL NUMBER - PAPER COPY
D2057-3-13
MANUAL NUMBER - AEROFICHE
D2057-3-13AF
TEMPORARY REVISION NUMBER
D2057-3TR8
MANUAL DATE
REVISION NUMBER
10 September 1982
3
DATE
1 March 1996
This Temporary Revision consists of the following pages, which affect and replace existing pages
in the paper copy manual and supersede aerofiche information.
SECTION
2
2
PAGE
27
32
AEROFICHE
FICHE/FRAME
1/B22
1/C03
SECTION
PAGE
AEROFICHE
FICHE/FRAME
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at the appropriate locations and remove and discard the superseded pages.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information on the frame (page) is no longer valid and the Temporary Revision should be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
cessna
A Textron Company
TEMPORARY REVISION NUMBER 7
DATE 7 October 2002
MANUAL TITLE
Model 210 & T210 Series 1977 Thru 1984 Service Manual
MANUAL NUMBER - PAPER COPY
D2057-3-13
MANUAL NUMBER - AEROFICHE
D2057-3-13AF
TEMPORARY REVISION NUMBER
D2057-3TR7
MANUAL DATE
10 September 1982
REVISION NUMBER
3
DATE
1 March 1996
This Temporary Revision consists of the following pages, which affect and replace existing pages
in the paper copy manual and supersede aerofiche information.
SECTION
2
2
2
2
2
2
2
2
2
2
2
16
16
PAGE
28
28A/Deleted
29
30
31
32
32A/Deleted
33
34
35
36
22C
22D
AEROFICHE
FICHE/FRAME
1/B23
NA
1/B24
1/C01
1/C02
1/C03
NA
Added
Added
Added
Added
Added
Added
SECTION
PAGE
AEROFICHE
FICHE/FRAME
REASON FOR TEMPORARY REVISION
1.
To include the requirement to inspect all fluid carrying lines and hoses in the cabin and wing areas.
Revise the Special Inspection Items section and 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 @ 2002
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA
TEMPORARY REVISION NUMBER 6
DATED 7 January 2000
MANUAL TITLE
MODEL 210 & T210 SERIES 1977 THRU 1984 SERVICE MANUAL
MANUAL NUMBER - PAPER COPY D2057-3-13
TEMPORARY REVISION NUMBER PAPER COPY D2057-3TR6
MANUAL DATE
D2057-3-13AF
AEROFICHE
10 SEPTEMBER 1982 REVISION NUMBER
3
AEROFICHE
DATE
N/A
1 MARCH 1996
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
28A
32A
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 a 2000
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA
TEMPORARY REVISION NUMBER 5
DATED
MANUAL TITLE
MODEL 210 SERIES 1977 THRU 1984 SERVICE MANUAL
MANUAL NUMBER - PAPER COPY
AEROFICHE
D2057-3-13
TEMPORARY REVISION NUMBER - PAPER COPY
MANUAL DATE
2 March, 1998
10 September, 1982
D2057-3TR5-13
REVISION NUMBER
3
D2057-3-13AF
AEROFICHE
DATE
N/A
1 March, 1996
This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy
manual and supersede aerofiche information.
CHAPTER/
SECTION/
SUBJECT
2
2
2
PAGE
AEROFICHE
FICHE/FRAME
30
31
32
CHAPTER/
SECTION/
SUBJECT
PAGE
AEROFICHE
FICHE/FRAME
1 C-01
1 C-02
1 C-03
REASON FOR TEMPORARY REVISION
To add Parker Hannifin Vacuum Manifold Check Valve inspection/replacement times to inspection section.
FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
For Paper Publications:
File this cover sheet behind the publication's title page to identify inclusion of the temporary revision in the
manual. Insert the new pages in the publication at the appropriate locations and remove and discard the
superseded pages.
For Aerofiche Publications:
Draw a line, with a 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 © 1998
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA
TEMPORARY REVISION NUMBER 4
DATED
MANUAL TITLE
Model 210, And T210 Series
1977 Thru 1984 Service Manual
MANUAL NUMBER - PAPER COPY
D2057-3-13
TEMPORARY REVISION NUMBER - PAPER COPY
MANUAL DATE
October 1, 1997
10 September 1982
AEROFICHE
D2057-3TR4-13
REVISION NUMBER
3
D2057-3-13AF
AEROFICHE
DATE
N/A
1 March 1996
This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy
manual and supersede aerofiche information.
CHAPTER/
SECTION/
SUBJECT
PAGE
AEROFICHE
FICHE/FRAME
1
5
1
6
1A15
1
7
Added
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
14
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
3
Added
Added
Added
Added
Added
Added
Added
Added
Added
Added
Added
Added
Added
Added
Added
2H10
CHAPTER/
SECTION/
SUBJECT
PAGE
AEROFICHE
FICHE/FRAME
1A14
REASON FOR TEMPORARY REVISION
1. To add wet torque values for McCauley propeller hub bolts and add standard torque value tables.
FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION
For Paper Publications:
File this cover sheet behind the publication's title page to identify inclusion of the temporary revision in the
manual. Insert the new pages in the publication at the appropriate locations and remove and discard the
superseded pages.
For Aerofiche Publications:
Draw a line, with a 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 © 1997
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA
LIST OF EFFECTIVE PAGE1
INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES.
NOTE: The portion of the text affected by the changes is indicated
by a vertical line in the outer margins of the page.
Changes to illustrations are indicated by miniature pointing hands.
2
Dates of issue for original and revised pages are:
Original ......
0....... 10 September 1982
Revision ...... 1....... 3 October 1983
Revision ...... 2 ....... 29 November 1983
Revision ...... 3 ....... 1 March 1996
TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 802, CONSISTING OF THE FOLLOWING:
Revision
No.
Page
No.
*Title ......................
*AthruB ..................
C Blank ...................
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*10-1 ......................
10-2 thru 10-8 .............
*11-1 ......................
11-2 thru 11-3 ..............
11-4 .......................
Upon receipt of the second and subsequent revisions to this book, personnel
responsible for maintaining this publication in current status should ascertain
that all previous revisions have been received and incorporated.
* The asterisk indicates pages revised, added, or deleted by the current revision.
A
Revision 3
Revision
No.
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LIST OF EFFECTIVE PAGES, Cont. Page
No.
Revision
No.
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...
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..........
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* The asterisk indicates pages revised, added, or deleted by the current revision.
Revision3
B/(C blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
TABLE OF CONTENTS
SECTION
PAGE NO.
AEROFICHE/MANUAL
1.
GENERAL DESCRIPTION ..........
2.
GROUND HANDLING, SERVICING. CLEANING,
.....
............
....
LUBRICATION AND INSPECTION ..........................
3.
FUSELAGE
4.
WINGS AND EMPENNAGE .............................
5.
LANDING GEAR, BRAKES AND HYDRAULIC SYSTEM
5A.
1A20/2-1
.........................................
(THRU 1978 MODELS)
1A10 1-1
... 1C9/3-1
1...
D20/4-1
..................................
1E5/-1
LANDING GEAR, BRAKES AND HYDRAULIC SYSTEM
(BEGINNING WITH 1979 MODELS) .........
1......1115/5A-1
6.
AILERON CONTROL SYSTEM
.............................
1K16/6-1
7.
WING FLAP CONTROL SYSTEM ...........................
1L3/7-1
8.
ELEVATOR CONTROL SYSTEM . ..........................
2A2/8-1
9.
ELEVATOR TRIM TAB CONTROL SYSTEM .....................
2A17/9-1
10.
RUDDER CONTROL SYSTEM ...............................
2B13/10-1
11.
RUDDER TRIM CONTROL SYSTEM ..........................
2C1/11-1
12.
ENGINE (NORMALLY ASPIRATED) ..........................
2C13/12-1
12A.
ENGINE (TURBOCHARGED)
2E6/12A-1
13.
FUEL SYSTEM ........................
14.
PROPELLERS AND PROPELLER GOVERNORS
15.
UTILITY SYSTEMS ...................................
2H16/15-1
16.
INSTRUMENTS AND INSTRUMENT SYSTEMS ..................
2K1/16-1
17.
ELECTRICAL SYSTEMS ...................................
3A2/17-1
18.
STRUCTURAL REPAIR ....................................
3D11/18-1
19.
EXTERIOR PAINTING ....................................
3E21/19-1
20.
WIRING DIAGRAMS ..................
............................
....
.................
|
2F19/13-1
2H6/14-1
...........
3F5/20-1
WARNING
When performing any inspection or maintenance that
requires turning on the master switch, installing a battery, or
pulling the propeller through by hand, treat the propeller as if
the ignition switch were ON. Do not stand nor allow anyone
else to stand, within the arc of the propeller, since a loose or
broken wire or a component malfunction could cause the
propeller to rotate.
Revision 3
i
MODEL 210 & T210 SERIES SERVICE MANUAL
CROSS REFERENCE LISTING
OF POPULAR NAME VS. MODEL NUMBERS AND SERIALS
All aircraft, regardless of manufacturer, are certified under model
number designations. However, popular names are often used for
marketing purposes. To provide a consistent method of referring
to these aircraft, the model number will be used in this publication
unless the popular name is necessary to differentiate between versions of the same basic model. The following table provides a listing of popular name, model number and serial number.
ii
SERIAL
POPULAR NAME
MODEL
YEAR
MODEL
BEGINNING
ENDING
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION II
1977
1977
1977
1977
210M
T210M
210M
T210M
21061574
21061574
21061574
21061574
21062273
21062273
21062273
21062273
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION
1978
1978
1978
1978
210M
T210M
210M
T210M
21062274
21062274
21062274
21062274
21062954
21062954
21062954
21062954
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION I
1979
1979
1979
1979
210M
T210M
210M
T210M
21062955
21062955
21062955
21062955
21063640
21063640
21063640
21063640
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION II
1980
1980
1980
1980
210M
T210M
210M
T210M
21063641
21063641
21063641
21063641
21064135
21064135
21064135
21064135
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION I
1981
1981
1981
1981
210N
T210N
210N
T210N
21064136
21064136
21064136
21064136
21064535
21064535
21064535
21064535
CENTURION
TURBO CENTURION
CENTURION I
TURBO CENTURION II
1982
1982
1982
1982
210N
T210N
210N
T210N
21064536
21064536
21064536
21064536
21064772
21064772
21064772
21064772
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION II
1983
1983
1983
1983
210N
T210N
210N
T210N
21064773
21064773
21064773
21064773
21064822
21064822
21064822
21064822
CENTURION
TURBO CENTURION
CENTURION II
TURBO CENTURION II
1984
1984
1984
1984
210N
T210N
210N
T210N
21064823
21064823
21064823
21064823
21064897
21064897
21064897
21064897
Revision 3
MODEL 210 &T210 SERIES SERVICE MANUAL
INTRODUCTION
This manual contains factory-recommended procedures and instructions for ground handling, servicing, and
maintaining Cessna 210 Series Models. The 210 and T210 Series Models covered in this manual are
identical, except the Model T210 is turbocharged. Besides serving as a reference for the experienced
mechanic, this book also covers step-by-step procedures for the less experienced.
This service manual is designed for aerofiche presentation. To facilitate the use of the aerofiche, refer to the
aerofiche header for basic information.
IMPORTANT INFORMATION CONCERNING
KEEPING CESSNA PUBLICATIONS CURRENT
The information in this publication is based on data available at the time of publication and is updated,
supplemented, and automatically amended by all information issued in Service News Letters, Service
Bulletins, Supplier Service Notices, Publication Changes, Revisions, Reissues and Temporary Revisions. All
such amendments become part of and are specifically incorporated within this publication. Users are urged
to keep abreast of the latest amendments to this publication through the Cessna Product Support
subscription services. Cessna Service Stations have also been supplied with a group of supplier publications
which provide disassembly, overhaul, and parts breakdowns for some of the various supplier equipment
items. Suppliers publications are updated, supplemented, and specifically amended by supplier issued
revisions and service information which may be reissued by Cessna; thereby automatically amending this
publication and is communicated to the field through Cessna's Authorized Service Stations and/or through
Cessna's subscription services.
IWARNING
ALL INSPECTION INTERVALS, REPLACEMENT TIME LIMITS, OVERHAUL
TIME LIMITS, THE METHOD OF INSPECTION, LIFE LIMITS, CYCLE LIMITS,
ETC., RECOMMENDED BY CESSNA ARE SOLELY BASED ON THE USE OF
NEW, REMANUFACTURED, OR OVERHAULED CESSNA APPROVED PARTS.
IF PARTS ARE DESIGNED, MANUFACTURED, REMANUFACTURED,
OVERHAULED, PURCHASED, AND/OR APPROVED BY ENTITIES OTHER
THAN CESSNA, THEN THE DATA IN CESSNA'S MAINTENANCE/SERVICE
MANUALS AND PARTS CATALOGS ARE NO LONGER APPLICABLE AND THE
PURCHASER IS WARNED NOT TO RELY ON SUCH DATA FOR NON-CESSNA
PARTS. ALL INSPECTION INTERVALS, REPLACEMENT TIME LIMITS,
OVERHAUL TIME LIMITS, THE METHOD OF INSPECTION, LIFE LIMITS,
CYCLE LIMITS, ETC., FOR SUCH NON-CESSNA PARTS MUST BE OBTAINED
FROM THE MANUFACTURER AND/OR SELLER OF SUCH NON-CESSNA
PARTS.
1. REVISIONS/CHANGES. Revisions/changes are issued for this publication as required and
include only pages that require updating.
2. REISSUE. A reissue is issued as required, and is a complete manual incorporating all the latest |
information and outstanding revisions/changes. It supersedes and replaces previous issue(s).
REVISIONS/CHANGES and REISSUES can be purchased from a Cessna Service Station or directly from
Cessna Parts Distribution (CPD 2), Dept. 701, Cessna Aircraft Company, P. O. Box 949, Wichita, Kansas
67201 (walk-in address: 5800 East Pawnee, Wichita, Kansas 67218).
All supplemental service information concerning this manual is supplied to all appropriate Cessna Service
Stations so that they have the latest authoritative recommendations for servicing these Cessna airplanes.
Therefore, it is recommended that Cessna owners utilize the knowledge and experience of the factory-trained
Service Station Organization.
Revision 3
iii
MODEL 210 & T210 SERIES SERVICE MANUAL
CUSTOMER CARE SUPPLIES AND PUBLICATIONS CATALOG
A Customer Care Supplies and Publications Catalog is available from a Cessna Service Station or directly
from Cessna Parts Distribution (CPD 2), Dept. 701, Cessna Aircraft Company, P. 0. Box 949, Wichita,
Kansas 67201 (walk-in address: 5800 East Pawnee, Wichita, Kansas 67218). This catalog lists all
publications and Customer Care Supplies available from Cessna for prior year models as well as new
products. To maintain this catalog in a current status, it is revised quarterly and issued on Aerofiche with
the quarterly Service Information Summaries. A listing of all available publications is issued periodically
by the Cessna Propeller Product Support Department.
SUPPLEMENTAL TYPE CERTIFICATE INSTALLATIONS
Inspection, maintenance, and parts requirements for supplemental type certificate (STC) installations are
not included in this manual. When an STC installation is incorporated on the airplane, those portions of the
airplane affected by the installation must be inspected in accordance with the inspection program published
by the owner of the STC. Since STC installations may change systems interface, operating characteristics,
and component loads or stresses on adjacent structures, Cessna provided inspection criteria may not be valid
for airplanes with STC installations.
CUSTOMER COMMENTS ON MANUAL
Cessna Aircraft Company has endeavored to furnish you with an accurate, useful, up-to-date manual. This
manual can be improved with your help. Please use the return card, provided with your manual, to report
any errors, discrepancies, and omissions in this manual as well as any general comments you wish to make.
iv
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 1
GENERAL DESCRIPTION
Page No.
Aerofiche/Manual
GENERAL DESCRIPTION .......
Model 210 Series . .......
Description .........
1A10/1-1
1A10/1-1
1A10/1-1
Aircraft Specifications
Stations ..........
Bolt Torques. .......
.. A10/1-1
1A10/1-1
1A14/1-5
1-3. GENERAL DESCRIPTION.
1-2.
MODEL 210-SERIES.
1-3. DESCRIPTION. The Cessna Centurion,
Centurion II, Turbo Centurion, and Turbo
Centurion II (Model 210 Series) aircraft, described
in this manual, are single-engine, high-wing
monoplanes of all metal, semimonocoque construction. Wings are full cantilever, with sealed sections
forming fuel bays. The fully-retractable tricycle
landing gear consists of tublar spring-steel main
gear struts and a steerable nose gear with an airhydraulic fluid shock strut. The six place seating
arrangement is of conventional, forward facing type.
Powering the Model 210 Series is a Continental, horizontally-opposed, air-cooled, six-cylinder, fuelinjected engine driving an all-metal, constant-speed
propeller. A more desirable higher performance
aircraft, is offered in the turbocharged version of the
Model 210 Series.
1-4. 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 on 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.
1-5. STATIONS. A station diagram is shown in figure 1-2 to assist in locating equipment when a written
description is inadequate or impractical
Revision 2
1-1
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 AND T210 SERIES
MAXIMUM WEIGHT - 210
Ramp ............................
3812 lbs.
Takeoff or Landing ......................
3800 Ibs.
STANDARD EMPTY WEIGHT - 210
Centurion ..........................
. 2173 lbs.
Centurion II .....................
.. 2223 lbs.
MAXIMUM USEFUL LOAD - 210
Centurion .........................
1639 lbs.
Centurion II . . . . . . . . . . . . . . . . . . . . . . . . . 1589 bs.
MAXIMUM WEIGHT - T210
Ramp ...........
..........
4016 lbs.
Takeoff ..........................
. 4000 lbs.
Landing ..................
........
3800 lbs.
STANDARD EMPTY WEIGHT - T210
Turbo Centurion ......................
. 2263 lbs.
Turbo Centurion II ......................
2311 lbs.
MAXIMUM USEFUL LOAD - T210
Turbo Centurion .......................
1753 lbs.
Turbo Centurion II ......................
1705 lbs.
FUEL CAPACITY
Total .................
........
...
90 gal.
Usable - Thru Serial 21064535 . . ..............
.89 gal.
Usable - Beginning with Serial 21064536 ...........
87 gal
OIL CAPACITY ................
.........
10 qt.
With External Oil Filter and
All Turbocharged Engines
................
11 qt.
ENGINE MODEL
210 (Refer to Section 121for Engine Data) ..........
T210 (Refer to Section 12A for Engine Data) ...........
PROPELLER (Constant-Speed)
(Three Blades) ...................
....
LANDING GEAR (Retractable, Hydraulically-Actuated) ........
MAIN WHEEL TIRES .......................
Pressure ..
....
. . . .......
. ..
. ....
NOSE WHEEL TIRE
210
........
..............
Pressure
..............
T210 (THRU T21062954) ...................
Pressure
.........
.
.............
T210 (BEGINNING WITH T21062955)
.............
Pressure
.
.
.
.
........
Figure 1-1.
1-2
Revision 2
........
. ....
. CONTINENTAL 10-520
CONTINENTAL TSIO-520
80" McCAULEY
Tricycle
6.00 x 6
. 55 psi
5.00 x 5, 6ply
50 psi
5.00 x 5, 6 ply
50 psi
5.00 x 5, 10 ply
88 psi
Aircraft Specifications (Sheet 1 of 2)
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 AND T210 SERIES
90 psi
NOSE GEAR STRUT PRESSURE (Strut Extended) ..........
WHEEL ALIGNMENT
Camber .
.........
Toe-in ..................
...........
4°
1° 30'
0" to . 06"
.....
.....
...
AILERON TRAVEL
Up
..............
......
WING FLAP TRAVEL (Electrically-Actuated)
. . . . . . . . . ..
. . . . . ..
°
0° ±0° to 30° , +1° -2 °
24 ° ± 1°
24°± 1°
. . . . .
27°13' ± 1 °
27° 13' ± 1°
23 ° ± 1°
. . . . . .
ELEVATOR TRIM TAB TRAVEL
Up
20° ±2
°
15°±2
............
RUDDER TRAVEL (Measured parallel to water line)
Right ......................
Left.
RUDDER TRAVEL (Measured perpendicular to hinge line)
....
. ....
. . . . . ...
....
Right ..
Left
............................
ELEVATOR TRAVEL
Up .............................
Down . . ..
.
....
Down ...................
. . . . . . . . . . . . . . . . . . . . . . . . . . . ..
. . . . . . . . . . . . .
Down . . . . . . . . . . . . . ..
PRINCIPAL DIMENSIONS
. . .
Wing Span .......................
..
. ....
. . . . .
. . . ..
.....
Tail Span .. ..
........
Length ..................
Fin Height (Maximum with Nose Gear Depressed and
Flashing Beacon Installed on Fin) ..............
Track Width .................
. ......
BATTERY LOCATION ....................
17° ± 1°
25 ° ±1°
10 °± 1°
441.75"
156.32"
337.96"
112.92"
104.20"
Left Side of Firewall
Figure 1-1. Aircraft Specifications (Sheet 2 of 2)
Revision 2
1-3
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 SERIES
25.25
54.00
~~°°11.
00
* INDICATES7796.00
CANTED BULKHEAD
138. 00
44.0 55.668.00 *106.0
0.0
3.8
70.8
*124.6
112.0
152.2
209.0
7
209.0
206.00
FureReference
1-2.
Stations0
18.0I
~~~1-4 Revis~~i~~on
3.8
8.1
0.0
2~~189.00
70.8
112,0
_90o.o
/
138.0152.2 180. 6
67.2
98.0
44 0 55.6
109.6*
3 106.0
166.4
'124.6
77.0
*INDICATES
CANTED BULKHEAD
Figure 1-2.
1-4
Revision 2
Reference Stations
194.8
23o. 8
MODEL 210 & T210 SERIES SERVICE MANUAL
1-6.
MATERIALAND TOOL CAUTIONS- GENERAL
A. Mercury
CAUTION
TEST
OTHER
AND
THERMOMETERS
EQUIPMENT CONTAINING MERCURY, MUST
NOT BE USED ON THE AIRPLANE.
Mercury, by the amalgamation process, can penetrate
any break in the finish, paint or sealing coating of a
metal structural element. An oxide coating on a dry
metallic surface will tend to inhibit an immediate
action while a bright, polished, shining or scratched
surface will hasten the process. Moisture will also
promote the amalgamation process. Soils, greases or
other inert contaminants, present on the metal
surfaces, will prevent the start of the action. The
corrosion and embrittlement which results from an
initial penetration, can be extremely rapid in
structural members under load. Once it has begun,
there is no known method of stopping it. Complete
destruction of the load carrying capacity of the metal
will result.
b. Maintenance Precautions
WARNING
DURING
MAINTENANCE,
REPAIR
AND
SERVICING OF THE AIRPLANE,
MANY
SUBSTANCES
AND
ENVIRONMENTS
ENCOUNTERED MAY CAUSE INJURY IF
PROPER
PRECAUTIONS
ARE
NOT
OBSERVED.
Carefully read and follow all instructions, and
especially adhere to all cautions and warnings
provided by the manufacturer of the product being
used. Use appropriate safety equipment as required
including goggles, face shields, breathing apparatus,
protective clothing and gloves. Fuel, engine oil,
solvents, volatile chemicals, adhesives, paints and
strong cleaning agents may cause injury when
contacting the skin or eyes, or when vapors are
breathed. When sanding composites or metals or
otherwise working in an area where dust particles
may be produced, the area should be ventilated and
the appropriate respirator must be used.
Solvents are hazardous to work with because of their
flammability, rate of evaporation and reaction to
oxidizers. Solvents can also be an irritant to the skin
and eyes.
A single spark, a smoldering cigarette, or even
atmospheric conditions can ignite solvent vapors. The
lower the flash point of the chemical, the more likely it
is to become flammable. Generally, flash points of less
than 100°F (37.8°C) are considered flammables.
Examples of solvent flash points are shown below:
SOLVENT
FLASH-POINT
Methyl Propyl Ketone
45°F
(7.2° C)
Touluene
39°F
(3.9° C)
Isopropyl Alcohol
53.6°F (12° C)
Acetone
1.4°
(-17°C)
The rate of evaporation is closely tied to flammability,
because normally the vapors must be present to ignite the
liquid. Vaporization also allows solvents, even those that
are not flammable, to get into the air and into the body's
blood stream through the lungs.
Solvents can also react explosively with oxidizers
(chemicals which release oxygen). A very violent and
uncontrollable reaction takes place which generates heat
rapidly. For this reason, it is very important for each
person to be aware of specific chemicals in use in the work
area, and to adhere to the labeling of containers. Chemical
manufacturers are required to label each container with a
diamond shaped symbol: red forflammable and yellow for
oxidizers.
Solvents can also damage the hands and skin. Solvents
dry out skin and dissolve the natural oils. The condition
can develop into an irritation, or if left untreated with
continuous exposure, it may progress to a dermatitis.
Damaged skin allows other contaminants to worsen the
condition, because the contaminants have easier access to
the deeper levels of the skin. In serious cases, blood
poisoning is also possible.
The best defense against skin irritation is not to be
exposed. If exposure is unavoidable, steps should betaken
to limit exposure times. Prolonged exposure to these
irritants can lead to long term liver damage.
c. General Usage Solvents
General usage solvents include the following:
Methyl Propyl Ketone
Toluene
Isopropyl Alcohol
Acetone
Methylene Chloride
1,1,1-Trichloroethane
Naptha
Trichloroethylene
These chemicals/solvents are generally colorless,
evaporate quicker than water, and tend to give off vapors
in higher quantities as their temperature increases. The
vapors are generally heavier than air, which causes them
to collect in low lying areas or push normal oxygen and
air out of a confined area. This situation can lead to
Temporary Revision Number 4
October 1,1997
1-5
MODEL 210 & T210 SERIES SERVICE MANUAL
1-7. TORQUE DATA- MAINTENANCE PRACTICES
f. Countersunk washers used with close tolerance
bolts must be installed correctly to ensure proper
torquing (refer to Figure 1-5).
To ensure security of installation and prevent over
stressing of components during installation, thetorque
applicable
this section
and
other
values outlined in chapters
values per Table
Table
nutsto torque
torque values
g. Tighten
Tighten accessible
accessible nuts
used~
during
of this
should
manual
be
or screws with
to
nutplates,
1-1. Screws attached
installation and repair of components
threads not listed in Table 201 should be tightened
installation andrepair of components.
firmly, but not to a specific torque value. Screws
The torque value tables, listed in this section, are
used with dimpled washers should not be drawn
standard torque values for the nut and bolt
tight enough to eliminate the washer crown.
combinations shown. If a component requires special
torque values, those values will be listed in the
applicable maintenance practices section
h. Table 1-1 is not applicable to bolts, nuts and screws
used in control systems or installations where the
.
Torque is typically applied and measured using a
required torque would cause binding or would
torquewrench. Differentadapters, used inconjunction
interfere with proper operation of parts. On these
with the torque wrench, may produce an actual torque
installations, the assembly should be firm but not
to the nut or bolt which is different from the torque
binding.
reading. Figure 1-4 is provided to help calculate actual
torque in relation to specific adaptors used with the
i. Castellated Nuts.
torquewrench
Free Running Torque Value
Free running torque value is the torque vale
required to rotate a nut on a threaded shaft,
tightening. Free running torque value does
without
not represent the torque values listed in the tables
of this section. Torque values listed in the tables
represent the torque values above free running
torque.
Self-locking and non self-locking castellated nuts,
tightened to the minimum torque value shown in
h 11 The torque may increased
nallth
Table 1-1. The torque may be increased to install the
cotter pin, but this increase must not exceed the
alternatetorque values.
MS17826 self-locking, castellated nuts shall be
torqued per Table 1-1.
EXAMPLE
If finaltorque required isto be 150 inch-pounds and
the free running torque is 25 inch-pounds, then the
free running torque must be added to the required
torque to achieve final torque of 150 +25 = 175
inch-pounds.
The end of the bolt or screw should extend through
the nut at least two full threads including the
chamfer.
Breakaway torque value is the value of torque
required to start a nut rotating on a thread shaft, and
does not represent free running torque value. It should
be noted that on some installations the breakaway
torque value cannot be measured.
General Torquing Notes:
a. These requirements do not apply to threaded parts
used for adjustment, such as turnbuckles and rod
ends.
b. Torque values shown are for clean, nonlubricated
parts. Threads should be free of dust, metal filings,
etc. Lubricants, other than that on the nut as
purchased, should not be used on any bolt
installation unless specified.
c. Assembly of threaded fasteners, such as bolts,
screws and nuts, should conform to torque values
shown in Table 1-1.
d. When necessary to tighten from the bolt head,
increase maximum torque value by an amount
equal to shank friction. Measure shank friction
with a torque wrench.
e. Sheet metal screws should be tightened firmly, but
notto a specifictorquevalue.
Temporary Revision Number 4
October 1, 1997
1-6
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE:
SHORT OPEN END
ADAPTER
WHEN USING A TORQUE WRENCH ADAPTER WHICH
CHANGES THE DISTANCE FROM THE TORQUE WRENCH
DRIVE TO THE ADAPTER DRIVE, APPLY THE FOLLOWING FORMULAS TO OBTAIN THE CORRECTED TORQUE
READING.
TORQUE
WRENCH
WRENCH
DRIVE
CENTERLINE
ADAPTER
DRIVE
CENTERLINE
HANDGRIP
CENTERLINE
(PREDETERMINED)
--
SETSCREWADAPTER
FORMULA TL
y
L+E
EXAMPLE (WITH "E" AS PLUS DIMENSION)
T
Y
E
L
HOSE CLAMP
ADAPTER
<
=
=
=
=
LEGEND
T = ACTUAL (DESIRED) TORQUE
=~Y
APPARENT(INDICATED) TORQUE
L = EFFECTIVE LENGTH LEVER
= EFFECTIVE LENGTH OF EXTENSION
)\~Qk)
<~~~--~~
OPEN-END WRENCH
ADAPTER
135x10 =117.39
10+1.5
Y = 117 IN-LB
y-
135IN-LB
UNKNOWN
1.5 IN
10.0 IN
~E
WRENCH
DRIVE
CENTERLINE-
ADAPTER
DRIVE
CENTERLINE
HANDGRIP
CENTERLINE
(PREDETERMINED)
FLARE NUT WRENCH
ADAPTER
TORQUE
FORMULA
~~~EXAMPLE (WITH "E" AS MINUS DIMENSION)
__ ^
SPANNER WRENCH
ADAPTER
WRENCH
T
Y
L
E
=
=
=
=
135 IN-LB
UNKNOWN
10.0 IN
1.5IN
y
135 x 10 = 1350 = 158.82
10 - .5
Y = 159 IN-LB
4^^.~~~~~h~~~~~~
ssss~5598C200
Torque Wrench and Adapter Formulas
Figure 1-4 Sheet 1
Temporary Revision Number 4
October 1,1997
1-7
MODEL 210 & T210 SERIES SERVICE MANUAL
EXTERNAL WRENCHING HEAD
CORRECT INSTALLATION
INSTALL WASHER WITH COUNTERSUNK
FACE NEXT TO BOLT HEAD RADIUS
INTERNAL WRENCHING HEAD
I
COUNTERSUNK
WASHER
STANDARD
WASHER
INCORRECT INSTALLATION
CAUTION:
NEVER INSTALL STANDARD
WASHER OR COUNTERSUNK
WASHER IN REVERSE WHEN
USING BOLTS WITH RADIUS
UNDER THE HEAD
5598C1004
5598C1004A
Washer Installation Close Tolerance Bolts
Figure 1-5 Sheet 1
Temporary Revision Number 4
October 1, 1997
1-8
MODEL 210 & T210 SERIES SERVICE MANUAL
Table 1-1: Torque Requirements For Steel Bolts, Screws and Nuts (Inch-Pounds)
*~~~~~~
~~~FINE
FINE THREADED SERIES
(TENSION TYPE NUTS)
SIZE
THREADED SERIES
(SHEAR TYPE NUTS
EXCEPT MS17826)
Standard
Torque
Alternate
Torque
Standard
Torque
Alternate
Torque
8-36
12to 15
--
7to9
--
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
20to25
50 to 70
100to 140
160to 190
450to500
480 to 690
800to 1000
1100to 1300
2300to2500
2500to3000
3700 to 4500
5000to 7000
9000to 11000
20to28
50 to 75
100to 150
160to 260
450to560
480 to 730
800to 1070
1100to 1600
2300to3350
2500to4650
3700 to 6650
5000 to 10000
9000to 16700
12to15
30 to 40
60to85
95to 110
270to300
290 to 410
480to600
660to 780
1300to 1500
1500to 1800
2200 to 3300
3000 to 4200
5400to6600
12to19
30 to 48
60to 100
95to 170
270to390
290 to 500
480to750
660to 1060
1300to2200
1500to2900
2200 to 4400
3000 to 6300
5400to 10000
MS17826 NUTS
Standard
Torque
Alternate
Torque
12to15
30 to 40
60to80
95to 110
180to210
240 to 280
320to370
480to 550
880to1010
1500to 1750
2200 to 2700
3200 to 4200
5900to6400
12to20
30 to 45
60to 90
95to 125
180to 225
240 to 300
320to400
480to 600
880to 1100
1500to 1900
2200 to 3000
3200to 5000
5900to7000
Fine Thread Tension application nuts include: AN310, AN315, AN345, MS17825, MS20365, MS21044 through MS21048,
MS21078, NAS679, NAS1291.
FineThread Shearapplication nuts include: AN316, AN320, MS21025, MS21042, MS21043, MS21083, MS21245, NAS1022,
S1117.
Coarse Thread application nuts include: AN340, MS20341, MS20365, MS35649
Table 1-1: Torque Values (Newton Meters) Nuts, Bolts and Screws (Steel)
SIZE OF
BOLT,
NUTOR
SCREW
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
FINE THREADED SERIES
(TENSION TYPE NUTS)
Standard
Torque
1.4to 1.7
2.3to 2.8
5.6 to 7.9
11.3to 15.8
18.1 to 21.5
50.8 to 56-5
54.2 to 78.0
90.4to 113.0
124.3 to 146.9
259.9to 282.5
282.5to 339.0
418.0 to 508.4
564.9 to 790.9
1016.9to 1242.8
Temporary Revision Number 4
October 1,1997
Alternate
Torque
-2.3to 3.2
5.6to8.5
11.3to 16.9
18.1 to 29.4
50.8 to 63.3
54.2 to 82.5
90.4to 120.9
124.3 to 180.8
259.9to 378.5
282.5 to 525.4
418.0 to 751.3
564.9 to 1129.9
1016.9to 1886.9
FINE THREADED SERIES
(SHEAR TYPE NUTS
EXCEPT MS17826)
Standard
Torque
0.8 to 1.01.4to 1.7
3.4 to4.5
6.8to 9.6
10.7 to 12.4
30.5 to 33.9
32.8 to 46.3
54.2 to 67.8
74.6 to 88.1
146.9 to 169.5
169.5 to 203.4
248.6 to 372.9
339.0 to 474.5
610.1 to745.7
MS17826 NUTS
Alternate
Torque
Standard
Torque
Alternate
Torque
1.4to 2.1
3.4to 5.4
6.8to 11.3
10.7 to 19.2
30.5 to 44.1
32.8 to 56.5
54.2 to 84.7
74.6to 119.8
146.9 to 248.6
169.5to 327.7
248.6 to 497.1
339.0 to 711.8
610.1 to 1129.9
1.4to 1.7
3.4 to4.5
6.8 to 9.0
10.7 to 12.4
20.3 to 23.7
27.1 to 31.6
36.2 to 41.8
54.2 to 62.1
99.4to 114.1
169.5to 197.7
248.6 to 305.1
361.6to 474.5
666.6to723.1
1.4to 2.3
3.4to 5.1
6.8to 10.2
10.7 to 14.1
20.3 to 25.4
27.1 to 33.9
36.2 to 45.2
54.2 to 67.8
99.4to 124.3
169.5to 214.7
248.6 to 339.0
361.6 to 564.9
666.6to790.9
1-9
MODEL 210 & T210 SERIES SERVICE MANUAL
Torque Requirements for Hi-Lok Fasteners
Use Table 1-2 to determine torque requirements for Hi-Lok fasteners.
NOTE:
Thistable is used in conjunction with MS21042 self-locking nuts.
Table 1-2. Torque Values Hi-Lok Fasteners (Used with MS21042 Self-Locking Nuts)
NOMINAL
FASTENER
DIAMETER
ALLOY STEEL
180- 200 KSI
(INCH POUNDS)
ALLOY STEEL
180- 200 KSI
(NEWTON METERS)
6-32
8-32
10-32
1/4-28
5/16-24
3/8-24
7/16-20
1/2-20
8to 10
12to 15
20to 25
50 to 70
100 to 140
160to 190
450 to 500
480 to 690
0.9to 1.1
1.4to 1.7
2.3to 2.8
5.6 to 7.9
11.3 to 15.8
18.1 to 21.5
50.8 to 56.5
54.2 to 78.0
Torque Requirements for Electrical Current Carrying And Airframe Ground Fasteners
Use Table 1-3 to determine torque requirements for threaded electrical current carrying fasteners.
Torque values shown are clean, nonlubricated parts. Threads shall be free of dust and metal filings. Lubricants, other than on
the nut as purchased, shall not be used on any bolt installations unless specified in the applicable chapters of this manual.
All threaded electrical current carrying fasteners for relay terminals, shunt terminals, fuse limiter mount block terminals
and bus bar attaching hardware shall be torqued per Table 1-3.
NOTE:
There isno satisfactory method of determining the torque previously applied to a threaded fastener. When
retorquing, always back off approximately 1/4 turn or more before reapplying torque.
Use Table 1-4to determine torque requirements for threaded fasteners used as airframe electrical ground terminals.
Table 1-3. Torque Values Electrical Current Carrying Fasteners
FASTENER
DIAMETER
TORQUE VALUE
(INCH POUNDS)
TORQUE VALUE
(NEWTON METERS)
6-32
8-32
10-32
3/16
1/4
5/16
3/8
1/2
8to 12
13to 17
20 to 30
20 to30
40 to 60
80to 100
105to 125
130to 150
0.9to 1.4
1.5to 1.9
2.3 to 3.4
2.3 to3.4
4.5to 6.8
9.0 to 11.3
11.9to 14.1
14.7 to 16.9
Temporary Revision Number 4
October 1, 1997
1-10
MODEL 210 & T210 SERIES SERVICE MANUAL
Table 1-4. Torque Values Airframe Electrical Ground Terminals
FASTENER
DIAMETER
TORQUE VALUE
(INCH POUNDS)
TORQUE VALUE
(NEWTON METERS)
5/16
3/8
130 to 150
160 to 190
14.7 to 16.9
18.1 to 21.5
Torque Requirements for Rigid Tubing and Hoses
Use Table 1-5 to determine torque requirements fortubes and hoses.
Table 1-5. Tubing/HoseTorque Limits (Inch-Pounds)
Flared or Flareless Fitting with
Aluminum or Annealed Stainless Steel
Tubing, and Hose with Aluminum Inserts
Flared or Flareless Fitting with Steel
Tubing, and Hose with Steel Inserts
Hose
Size
Tubing
O.D.
-2
1/8
Min
45
Max
55
Min
65
Max
75
-3
-4
-5
-6
-8
-10
-12
-16
3/16
1/4
5/16
3/8
1/2
5/8
3/4
1
75
105
135
160
265
340
425
710
85
115
145
175
290
375
470
785
95
135
180
260
475
665
855
1140
105
150
200
285
525
735
945
1260
Table 1-5. Tubing/HoseTorque Limits (Newton Meters)
Hose
Size
Tubing
O.D.
-2
-3
-4
-5
-6
-8
-10
-12
-16
1/8
3/16
1/4
5/16
3/8
1/2
5/8
3/4
1
Temporary Revision Number4
October 1, 1997
Flared or Flareless Fitting with
Aluminum or Annealed Stainless Steel
Tubing, and Hose with Aluminum Inserts
Flared or Flareless Fitting with Steel
Tubing, and Hose with Steel Inserts
Min
Max
Min
Max
5.1
8.5
11.5
15.3
18.1
29.9
38.4
48.0
80.2
6.2
9.6
13.0
16.4
19.8
32.8
42.4
53.1
88.7
7.3
10.7
15.3
20.3
29.4
53.7
75.1
96.6
128.8
8.5
11.9
16.9
22.6
32.2
59.3
83.0
106.8
142.4
1-11
MODEL 210 &T210 SERIES SERVICE MANUAL
1-8.
SAFETYING - MAINTENANCE PRACTICES
Safety Wire Installation (Refer to Figure 1-6).
Safety Wire Inconel (Uncoated), Monel (Uncoated).
Used for general safety wiring purposes. Safety wiring
is the application of wireto prevent relative movement
of structural or other critical components subjected to
vibration, tension, torque, etc. Monel to be used at
temperatures up to 700°F (370ºC) and inconel to be
used at temperatures up to 1500 F (815°C). Identified
by the color of the finish, monel and inconel color is
natural wire color.
Copper, is cadmium plated and dyed yellow in accordance
with FED-STD 595.
This wire will be used for shear and seal wiring
applications. Shear applications are those where it is
CAUTION
CAUTION
SCREWS IN CLOSELY SPACED GEOMETRIC
PATTERNS WHICH SECURE HYDRAULIC OR
AIR SEALS, HOLD HYDRAULIC PRESSURE, OR
USED IN CRITICAL AREAS SHOULD USE THE
WIRING
CRITICAL
OF SAFETY
USED
DOUBLE TWIST METHOD OF SAFETY WIRING.
Single wire method of safety wiring shall use the largest
nominal size wire listed in Table 1-6, which will fit the
hole.
The double twist method of safety wiring shall be used as
or shear
break or
purposely break
necessary
to
wire to
the wire
shear the
to purposely
necessary to
permit operation or actuation of emergency devices.
Seal applications are those where the wire is used with
a lead seal to prevent tampering or use of a device
without indication. Identified bythe color of the finish,
copper wire is dyed yellow,
The double twist method of safety wiring shall be used as
the common method of safety wiring. It is really one wire
twisted on itself several times. The single wire method of
safety wiring may be used in a closely spaced, closed
geometrical pattern (triangle, square, circle, etc.), on
parts in electrical systems, and in places that would make
Aluminum Alloy (Alclad 5056), is anodized and dyed blue
in accordance with FED-STD 595.
the single wire method more advisable. Closely spaced
shall be considered a maximum of two inches between
centers.
This wire will be used exclusively for safety wiring
magnesium parts.
NOTE
Surface treatments which obscure visual identification
of safety wire is prohibited.
Use single wire method for shear and seal wiring
application. Make sure the wire is installed so that it can
be easily broken when required in an emergency
situation. For securing emergency devices where it is
necessary to break the wire quickly, use copper only.
Inconel or monel, wire can be substituted for same
diameter and length of carbon steel or corrosion resistant
wire.
Safety wiring by the double twist method shall be done as
follows:
Wires are visually identifiable by their colors: natural for
inconel and monel, yellow for copper, and blue for
aluminum.
One end of the safety wire shall be inserted through
one set of safety wire holes in the bolt head. The other
end of the safety wire shall preferably be looped firmly
around the head to the next set of safety wire holes in
the same unit and inserted through this set of safety
wire holes. The "otherend" may go overthe head when
the clearances around the head are obstructed by
adjacent parts.
Cotter Pin.
The selection of material shall be in accordance with
temperature, atmosphere and service limitations,
Safety Wire
The size of the safety wire shall be in accordance with
the requirements of Table 1-6.
0.032 inch diameter safety wire is for general
purpose use; however, 0.020 inch diameter safety
wire may be used on arts having a nominal hole
partsand
having
than between
of less
tediameter
nominal
0.062a
between 0.045
0.045
nominal hole
inch with spacing between parts of less than two
inches, or on closely spaced screws and bolts of 0.25
inch diameter and smaller.
The strands, while taut, shall be twisted until the
twisted part is just short of the nearest safety wire hole
in the next unit. The twisted portion shall be within 1/8
inch of the holes in each unit. The actual number of
twists will depend upon the wire diameter, with
smaller diameters being able to have more twists than
larger diameters. The twisting shall keep the wire taut
without over stressing or allowing it to become nicked,
kinked or mutilated. Abrasions from commercially
available twist pliers shall be acceptable.
0.020 inch diameter copper wire shall be used for
shear and seal wire applications.
When employing the single wire method of locking,
the largest nominal size wire for the applicable
material or part in which the hole will
accommodateshallbe used.
Temporary Revision Number 4
October 1, 1997
1-12
MODEL 210 & T210 SERIES SERVICE MANUAL
_BEND
~
STEP 1.
-
INSERT WIRE THROUGH BOLT A AND
AROUND BOLT (IF NECESSARY,
BEND WIRE ACROSS BOLT HEAD).
TWIST WIRES CLOCKWISE UNTIL
THEY REACH BOLT B.
STEP 2.
INSERT ONE END OF WIRE THROUGH
BOLT B. BEND OTHER END AROUND
BOLT (IF NECESSARY, BEND WIRE
ACROSS HEAD OF BOLT). TWIST
WIRES COUNTERCLOCKWISE 1/2 INCH
OR SIX TWISTS. CLIP ENDS.
BEND PIGTAIL BACK AGAINST PART.
NOTE:
RIGHT THREADED PARTS SHOWN:
REVERSE DIRECTIONS FOR LEFT PARTS.
BOLT B
CLOCKWISE
DOUBLE-WIRE SAFETYING
COUNTER-
CLOCKWISE
^^CLOCKWISE
^
^
COUNTERCLOCKWISE
^-CLOCKWISE
~
MULTIPLE FASTENER
APPLICATION DOUBLE
TWIST - MULTIPLE
HOLE METHOD.
DOUBLE-TWIST SAFETYING
SINGLE HOLE METHOD
5598C2001
5599C2001
6598C1029
Lockwire Safetying
Figure 1-6, Sheet 1
Temporary Revision Number 4
October 1,1997
1-13
MODEL 210 & T210 SERIES SERVICE MANUAL
EXTERNAL SNAP RING
SINGLE-WIRE METHOD
BOLTS IN CLOSELY SPACED, CLOSED
GEOMETRICAL PATTERN, SINGLE
WIRE METHOD
SINGLE FASTENER APPLICATION
DOUBLE-TWIST METHOD
SMALL SCREWS IN CLOSELY SPACED, CLOSED
GEOMETRICAL PATTERN, SINGLE WIRE METHOD
NOTE:
RIGHT THREADED
PARTS SHOWN. REVERSE
DIRECTION FOR LEFT
THREADS
_^^~~~~~~~~~~~
~~5598C1024
5598C1 003
5598C1024
5598C1024
Lockwire Safetying
Figure 1-6, Sheet 2
Temporary Revision Number 4
October 1, 1997
1-14
MODEL 210 & T210 SERIES SERVICE MANUAL
Lockwire Safetying
Figure 1-6, Sheet 3
Temporary Revision Number 4
October 1, 1997
1-15
MODEL 210 & T210 SERIES SERVICE MANUAL
Table 1-6. Safety Wire
MATERIAL
SIZE AND NUMBER (MS20995-XXX)
0.015
0.020
0.032
0.040
0.041
0.047
Ni-Cu Alloy
(Monel)
_
NC20
NC32
NC40
_
Ni-Cr-Fe Alloy
(Inconel)
_
N20
N32
N40
Carbon Steel
_
F20
F32
_
Corrosion Resistant
Steel
C15
C20
C32
_C41
Aluminum Alloy
(Blue)
_
AB20
AB32
_
Copper (Yellow)
CY15
CY20
The wire shall be twisted to form a pigtail of 3 to 5
twists after wiring the last unit. The excess wire shall
be cut off. The pigtail shall be bent toward the part to
prevent it from becoming a snag. Safety wiring
multiple groups by the double twist double hole
method shall be the same as the previous double twist
single hole method except the twist direction between
subsequent fasteners may
be clockwise
or
counterclockwise.
Spacing
AB41
0.091
NC51
NC91
N51
N91
F47
F91
C47
C91
AB47
_AB91
Usage
A pigtail of 0.25 to 0.50 inch (3 to 5 twists) shall be
made at the end of the wiring. This pigtail shall be bent
back or under to prevent it from becoming a snag.
Safety wire shall be new upon each application.
When castellated nuts are to be secured with safety
wire, tighten the nut to the low side of the selected
torque range, unless otherwise specified, and if
necessary, continue tightening until a slot aligns with
the hole.
When safety wiring widely spaced multiple groups by
the double twist method, three units shall be the
maximum number in a series.
When safety wiring closely spaced multiple groups,
the number of units that can be safety wired by a
twenty four inch length of wire shall be the maximum
number in a series.
Widely spaced multiple groups shall mean those in
which the fastenings are from four to six inches apart.
Safety wiring shall not be used to secure fasteners or
fittings which are spaced more than six inches apart,
unless the points are provided on adjacent parts to
shorten the span of the safety wire to less than six
inches.
Tension
Parts shall be safety wired in such a manner that the
safety wire shall be put in tension when the part tends
to loosen. The safety wire should always be installed
and twisted so that the loop around the head stays
down and does not tend to come up over the bolt head
and leave a slack loop.
NOTE
~~~~~~NOTE
~Drilled
This does not necessarily apply to castellated nuts
when the slot is close to the top of the nut, the wire
will be more secure if it is made to pass along the
side ofthe stud.
Care shall be exercised when installing safety wire to
ensure that it is tight but not over stressed.
Temporary Revision Number 4
October 1, 1997
F41
0.051
In blind tapped hole applications of bolts or castellated
nuts on studs, the safety wiring shall be as described in
these instructions.
Hollow head bolts are safetied in the manner
prescribed for regular bolts.
Drain plugs and pet cocks may be safetied to a bolt, nut
or other part having a free lock hole in accordance with
the instructions described in this text.
External snap rings may be locked, if necessary, in
accordance with the general locking principles as
described and illustrated. Internal snap rings shall not
be safety wired.
When safety wiring is required on electrical connectors
which use threaded coupling rings, or on plugs which
employ screws or rings to fasten the individualparts of
the plug together, they shall be safety wired with 0.020
inch diameter wire in accordance with the safety
wiring principles as described and illustrated. It is
preferable to safety wire all electrical connectors
individually. Do not safety wire one connector to
another unless it is necessary to do so.
head bolts and screws need not be safety wired
if installed into self-locking nuts or installed with lock
washers. Castellated nuts with cotter pins or safety
wire are preferred on bolts or studs with drilled shanks
but self-locking nuts are permissible within the
limitations of MS33588.
1-16
MODEL 210 & T210 SERIES SERVICE MANUAL
Larger assemblies, such as hydraulic cylinder heads
for which safety wiring is required but not specified,
shall be safety wired as described in these instructions.
Safetying Turnbuckles
Safetying Turnbuckles
Use of Safety Wire.
Some turnbuckles are secured using safety wire. These
Safety wire shall not be used to secure nor shall safety
safetying procedures are detailed and illustrated in
wire be dependent upon fracture as the basis for
Federal Publication AC 43-13.1A, Safety Methods For
operation of emergency devices such as handles,
switches, guards covering handles, etc., that operate turnbuckles.
emergency mechanism such as emergency exits, fire
Use of Locking Clips
extinguishers, emergency cabin pressure release,
emergency landing gear release and the like.
General instruction for the selection and application of
However, where existing structural equipment or
locking clips(RefertoFigures 1-8and 1-9).
safety of flight emergency devices require shear wire
Prior to safetying, both threaded terminals should be
to secure equipment while not in use, but which are
dependent upon shearing or breaking of the safety
screwed an equal distance into the turnbuckle barrel,
wireforsuccessful emergency operation of equipment,
and should be screwed in, at a minimum, so no more
particular care shall be exercised to that wiring under
than three threads of any terminal are exposed outside
the body
these circumstances shall not prevent emergency
operations of these devices.
After the turnbuckle has been adjusted to its locking
Cotter Pin Installation
position, with the groove on terminals and slot
General instruction for the selection and application of
indicator notch on barrel aligned, insert the end of the
cotter pins (Referto Figure 1-7).
locking clip into the terminal and barrel until the "U"
curved end of the locking clip is over the hole in the
Select cotter pin material in accordance with
center of the barrel.
temperature, atmosphere and service limitations.
Cotter pins shall be new upon each application.
a. Press the locking clip into the hole to its full extent.
When nuts areto be secured to thefastenerwith cotter
pins, tighten the nut to the low side (minimum) of the
applicable specified or selected torque range, unless
otherwise specified, and if necessary, continue
tightening until the slot aligns with the hole. In no
case shall the high side (maximum) torque range be
exceeded.
b. The curved end of the locking clip will latch in the
hole in the barrel.
Castellated nuts mounted on bolts may be safetied
withcotterpinsorsafetywire.Thepreferredmethodis
with the cotter pin. An alternate method where the
cotter pin is mounted normal to the axis of the bolt
may be used where the cotter pin in the preferred
method is apt to become a snag.
In the event of more than 50 percent of the cotter pin
diameter is above the nut castellation, a washer
should be used under the nut or a shorter fastener
should be used. A maximum of two washers may be
permitted under a nut.
c. To check proper seating of locking clip, attempt to
remove pressed "U" end from barrel hole with
fingers only.
NOTE
Do not use a tool as the locking clip could be
distorted.
Locking clips are for one time use only and should not
be reused.
Both locking clips may be inserted in the same hole of
the turnbuckle barrel or in opposite holes of the
turnbuckle barrel.
The largest nominal diameter cotter pin listed in
MS24665, which the hole and slots will accommodate,
shall be used; but in no application to a nut, bolt or
screw shall the pin size be less than the sizes described
in Figure 1-7.
Install the cotter pin with the head firmly in the slot of
the nut with the axis of the eye at right angles to the
bolt shank, and bend prongs so that the head and
upper prong are firmly seated against the bolt.
In the pin applications, install the cotter pin with the
axis of the eye parallel to the shank of the clevis pin or
rod end. Bend the prongs around the shank of the pin
or rod end.
Cadmium plated cotter pins shall not be used in
applications bringing them in contact with fuel,
hydraulic fluid or synthetic lubricants.
Temporary Revision Number 4
October 1, 1997
1-17
MODEL 210 & T210 SERIES SERVICE MANUAL
TO PROVIDE CLEARANCE
PRONG MAY BE CUT HERE
CASTELLATED NUT ON BOLT
ALTERNATE METHOD
CASTELLATED NUT ON BOLT
PREFERRED METHOD
THREAD SIZE
6
MINIMUM
PIN SIZE
(INCH)
0.028
8
0.044
10
1/4
0.044
0.044
5/16
0.044
3/8
0.072
7/16
0.072
1/2
0.072
9/16
9/16
5/8
0/086
0.086
0.086
3/4
0.086
7/8
0.086
1
0.086
11/8
0.116
1 1/4
0.116
1 3/8
0.116
1 1/2
0.116
TANGENT
TO PIN
MAXIMUM
COTTER PIN
LENGTH
\~
60 DEGREES
60 DEGREES
MINIMUM
COTTER PIN
LENGTH
PIN APPLICATION
_~~~~~~~~~~~
~5598C1025
5598C1025
5598C1025
5598C1025
Cotter Pin Safetying
Figure 1-7, Sheet 1
Temporary Revision Number 4
October 1,1997
1-18
MODEL 210 &T210 SERIES SERVICE MANUAL
STRAIGHT END
*
~HOOK
SHOULDER
END LOOP
HOOK LIP
HOOK LOOP
n
PULL FOR INSPECTION
PULL FOR INSPECTION
55982002
Safetying Tumbuckle Assemblies
Figure 1-8, Sheet 1
Temporary Revision Number 4
October 1, 1997
1-19
MODEL 210 & T210 SERIES SERVICE MANUAL
TURNBUCKLE
CLEVIS
LOCKING CLIP
MS21256
TURNBUCKLE EYE
CABLE
THIMBLE
TURNBUCKLE BARREL
MS21251
LOCKING CLIP
MS21 256
TYPICAL TURNBUCKLE ASSEMBLY
SWAGED
TERMINAL
METHOD OF ASSEMBLING LOCKING CLIPS, TURNBUCKLE BARREL AND TERMINALS
NOMINAL
CABLE DIA.
THREAD
UNF-3
LOCKING
CLIP
MS21256
(NOTE 1)
1/16
No. 6-40
-1
-2S
3/32
No. 10-32
-1
-3s
-2
-3L
-1
-4S
-2
-4L
-1
-5S
-2
-5L
-1
-6S
-2
-6L
-2
-7L
1/8
5/32
3/16
1/4-28
5/16-24
7/32
TURNBUCKLE
BODY
MS21251
1/4
3/8-24
-2
-8L
9/32
7/16-20
-3
-9L
5/16
1/2-20
-3
-10L
NOTE 1:
TWO LOCKING CLIPS REQUIRED FOR EACH
TURNBUCKLE.
5598C1023
5598C1023
Safetying Turnbuckle Assemblies
Figure 1-9, Sheet 1
Temporary Revision Number 4
October 1,1997
1-20
MODEL 210 & T210 SERIES SERVICE MANUAL
WIRE
BREAKAGE
1-9.
AND
CABLE
CONTROL
BREAKAGE
AND
CABLE
WIRE
CONTROL
CORROSION LIMITATIONS
. Cables.~
of. Control^
~Individual
Examination
Cables.of Control
Control cable assemblies are subject to a variety of
environmental conditions and forms of deterioration.
Some deterioration, such as wire or strand breakage, is
easy to recognize. Other deterioration, such as internal
corrosion or cable distortion, is harder to identify. The
following information will aid in detecting these cable
conditions.
Wire breakage criteria for cables in flap, aileron,
rudder, and elevator systems are as follows:
broken wires at random locations are
acceptable in primary and secondary control cables
when there are no more than six broken wires in
any given ten-inch cable length.
Corrosion
Broken Wire Examination (Referto Figure 1-9).
Examine cables for broken wires by passing a cloth
along length of cable. This will detect broken wires, if
cloth snags on cable. Critical areas for wire breakage
are those sections of cable which pass through
fairleads, across rub blocks, and around pulleys. If no
snags are found, then no further inspection is required.
If snags are found or broken wires are suspected, then
a more detailed inspection is necessary, which requires
that the cable be bent in a loop to confirm broken
wires. Loosen or remove cable to allow itto be bent in a
loop as shown. While rotating cable, inspect bent area
for broken wires.
remove and bend cable to properly inspect it for
internal strand corrosion, as this condition is usually
not evident on outer surface of cable. Replace cable if
internal corrosion is found. If a cable has been wiped
clean of its corrosion-preventive lubricant and metalbrightened, the cable shall be examined closely for
corrosion.
1-9.
Temporary Revision Number 4
October 1, 1997
Carefully examine any cable for corrosion that has a
broken wire in a section not in contact with wearproducing airframe components, such as pulleys,
1-21
MODEL 210 & T210 SERIES SERVICE MANUAL
BROKEN WIRE UNDETECTED BY
WIPING CLOTH ALONG CABLE
BROKEN WIRE DETECTED VISUALLY
WHEN CABLE WAS REMOVED
AND BENT
DO NOT BEND INTO LOOP SMALLER
THAN 50 CABLE DIAMETERS
NORMAL TECHNIQUE FOR
BENDING CABLE AND
CHECKING FOR BROKEN WIRES
Cable Broken Wire Examination
Figure 1-9 Sheet 1
Temporary Revision Number 4
October 1, 1997
1-22
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 2
GROUND HANDLING, SERVICING, CLEANING, LUBRICATION AND INSPECTION
WARNING
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the
propeller, since a loose or broken wire or a
component malfunction could cause the propeller
to rotate.
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
1A21/2-2
GROUND HANDLING ........
. 1A21/2-2
...........
Towing
1A21/2-2
.....
Hoisting ......
1A21/2-2
............
Jacking
. .. 1A21/2-2
Leveling ........
A21/2-2
Weighing ..........
. .. . 1A22/2-3
Parking ....
1A22/2-3
Tie-Down ..........
. 1A22/2-3
.
Flyable Storage ....
Returning Aircraft to Service . .. 1A22/2-3
. .1A22/2-3
Temporary Storage . .
. 1B1/2-6
Inspection During Storage ....
. . . B1/2-6
Returning Aircraft to Service
. .. 1B1/2-6
Indefinite Storage ...
1B2/2-7
Inspection During Storage .....
. . 1B2/2-7
Returning Aircraft to Service
1B3/2-8
.......
SERVICING .
1B3/2-8
.....
Description
1B3/2-8
....
Fuel Bays ...
Fuel Additives for Cold Weather
1B3/2-8
.....................
.
Operation
1B4/2-9
.............
Fuel Drains .
1B4/2-9
.................
Engine Oil ....
1B4/2-9
Engine Induction Air Filter .......
1B5/2-10
Vacuum System Air Filter ........
1B5/2-10
...............
Battery .........
1B6/2.11
.................
.
Tires
1B6/2-11
.............
Nose Gear Strut
Nose Gear Shimmy Dampener . . . 1B6/2-11
.1B6/2-11
Hydraulic Brake Systems ...
Landing Gear Hydraulic Retraction
1B6/2-11
.......
System
Hydraulic Fluid Sampling and
1B7/2-12
Contamination Check ......
1B7/2-12
. . . . .....
Oxygen System.
. 1.B7/2-12
Face Masks .........
1B7/2-12
. . . . . . . .
CLEANING ..
1B7/2-12
. ....
General Description ..
.. .1B7/2-12
Upholstery and Interior ...
...1B7/2-12
Plastic Trim .......
.1B7/2-12
.. .
Windshield and Windows.
. .1B7/2-12
Aluminum Surfaces . ..
1B7/2-12
Painted Surfaces .........
Engine and Engine Compartment . . 1B8/2-13
1B9/2-14
Propeller ........
1B9/2-14
Wheels ............
1B9/2-14.
......
LUBRICATION .
.1B9/2-14
General Description .....
1B9/2-14
Nose Gear Torque Links .....
.1B9/2-14
..
Tachometer Drive Shaft ..
Wheel Bearing Lubrication- . · . 1B9/2-14
1B9/2-14
Wing Flap Actuator........
1B9/2-14
.
..
Rod End Bearings ..
1B18/2-23
...........
INSPECTION
Revision 3
2-1
MODEL 210 & T210 SERIES SERVICE MANUAL
2-1.
GROUND HANDLING.
fuselage at the first bulkhead forward of the leading
edge of the stabilizer. If the optional hoisting rings
2-2. TOWING. Moving the aircraft by hand is accomplished by using the 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
CAUTIONWhen 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 hoisted with a
hoist of two-ton capacity, either by using hoisting
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
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 pro-
CAUTION I
When using the landing gear strut jack pad,
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
landing gear strut jack pad is not recommended
2-4A. LEVELING. Longitudinally leveling of the
aircraft is accomplished by backing out thetwo
screws on the left side of the fuselage and then
placing a level across the screws. Corresponding
points on either the upper or lower main door sills
may be used to level the aircraft laterally.
2-4B. WEIGHING AIRCRAFT.
Operating Handbook.
SHOP NOTES:
2-2
Refer to Pilot's
MODEL 210 & T210 SERIES SERVICE MANUAL
.
|
TOW BAR:
PART NUMBER 0501019-1, IS AVAILABLE FROM THE CESSNA SUPPLY DIVISION.
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 when they are
overheated or during cold weather when
accumulated moisture may freeze them.
a. Tie ropes, cables or chains to the wing tie-down
fittings located mid-wing in line with the outboard
edge of the flaps. 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 storage and/or the first 25 hours of intermittent engine
operation.
NOTE
Typical Tow Bar
Oil (Military Specification MIL-C-6529,
Type II). 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. In the event it is necessary to add
oil during the first 25 hours of operation use
only aviation grade straight mineral oil of the
correct viscosity.
During the 30 day non-operational storage or the first
25 hours of intermittent engine operation, every seventh day the propeller shall be rotated by hand without
running the engine. After rotating the engine five revolutions, stop the propeller 45º to 90* from the position
it was in. If the aircraft is stored outside, tie-down
in accordance with paragraph 2-8. In addition, the
pitot tube, static air vents, air vents, openings in the
engine cowling, and other similar openings shall have
protective covers installed to prevent entry of foreign
material. If at the end of thirty (30) days aircraft
will not be removed from storage, the engine shall
be started and run. The preferred method would be
to fly the aircraft for thirty (30) minutes, and up to,
but not exceeding normal oil and cylinder temperatures.
CAUTION
Excessive ground operation shall be avoided.
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 and clean oil pressure
screen (or 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
a maximum of 90 days. The aircraft is constructed
The aircraft is delivered from Cessna with
a Corrosion preventive Aircraft Engine
Revision 2
2-3
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
JACKING AIRCRAFT
1.
2.
3.
4.
5.
6.
7.
Lower the aircraft tail so that wing jack and stands can be placed at wing jack points.
Raise aircraft tail and attach tail stand to tail tie-down ring. BE SURE the tail stand
weighs enough to keep the tail down under all conditions and that it is strong enough to
support any weight that may be placed upon it.
Raise jacks evenly until desired height is reached. When jacking the aircraft, the main
landing gear wheels must be a minimum of 16" above shop floor for landing gear retraction.
The jack point on the bottom of the step may be used to raise only one main wheel.
Do not use brake casting as a jack point.
The nose may be raised by weighting down the tail. Place weight on each side of stabilizer,
next to fuselage.
Whenever the landing gear is to be operated in the shop, use the wing jack and tail jack points
to raise the aircraft.
The aircraft may be hoisted as outlined in paragraph 2-3.
REMOVING AIRCRAFT FROM JACKS
1.
2.
3.
4.
5.
6.
Place landin gear control handle in gear down position.
Operate ground hydraulic power source or aircraft emergency hydraulic hand pump until
landing gear is down and locked and the green indicator light is observed.
Disconnect ground hydraulic power source and/or stow emergency hydraulic hand pump handle.
Ascertain that green (DOWN) light is illuminated; then place master switch in OFF position.
Lower jacks evenly until aircraft rests on the landing gear and remove wing jacks and tail
stand.
Compress nose landing gear shock strut to static position.
SHOP NOTES:
Figure 2-2.
Jacking Details (Sheet 2 of 2)
2-5
MODEL 210 & T210 SERIES SERVICE MANUAL
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 bays 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 or oil.
d. Either block up fuselage to relieve pressure on
tires or rotate wheels every 30 days to 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 reinstalled 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 battery periodically and charge as required.
NOTE
An engine treated in accordance with the following may be considered being 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.
h. Using a portable pressure sprayer, spray preservative oil through the upper spark plug 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.
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.
2-6
1. Apply preservative oil to the engin 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. Remove dust collections as frequently as
possible. Clean and wax aircraft as required.
b. Inspect the interior of at least one cylinder
through the spark plug hole for corrosion at least
once each month.
NOTE
Do not move crankshaft when inspecting
interior of cylinder for corrosion.
c. If at the end of the 90 day period, the aircraft
is to be continued in non-operational storage, repeat
the procedural steps "g" thru "o" of paragraph 2-9.
AIRCRAFT SERVICE. After
RETURNING
2-11. RETURNING AIRCRAFT TO SERVICE. After
temporary storage, use the following procedure to
return the aircraft to service.
a. Remove aircraft from blocks. Check tires for
proper inflation.
b. Check and install battery.
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 values listed
in Section 12 or 12A of this manual.
i. Check fuel strainer. Remove and clean filter
screen, if necessary. Check fuel bays and fuel lines
for moisture and sediment. Drain enough fuel to
eliminate moisture and sediment.
Perform a thorough pre-flight inspection, then
j.
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 atmospheric corrosion, provided the
procedures outlined in paragraph 2-13 are performed
at the intervals specified.
MODEL 210 & T210 SERIES SERVICE MANUAL
a. Operate engine until oil temperature reaches
normal operating range. Drain engine oil sump and
reinstall & safety drain plug.
b. Fill oil sump to normal operating capacity with
corrosion preventive mixture which has been thoroughly mixed.
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.
c. Immediately after filling the oil sump with corrosion preventive mixture. fly the aircraft for a
period of time not to exceed a maximum of 30 minutes.
d. With engine operating at 1200 to 1500 rpm and
induction air filter removed, spray corrosion preventive mixt-re into induction airbox, at the rate of
one-half gallon per minute, until heavy smoke comes
from exhaust stack, then increase the spray until the
engine is stopped.
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
inside 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.
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.
p. Prepare airframe for storage as outlined in
paragraph 2-9 thru step "f."
CAUTION
NOTE
Injecting corrosion-preventive mixture too
fast can cause a hydrostatic lock.
As an altermate method of indefinite storage,
the aircraft may be serviced in accordance
with paragraph 2-9 providing the aircraft is
run up at maximun intervals of 90 days and
then reserviced per paragraph 2-9.
e. Do not rotate propeller after completing step
"d. "
f. Remove all spark plugs and spray corrosionpreventive mixture, which has been pre-heated
(221 ° to 2500F,) into all spark plug holes to thoroughly cover interior surfaces of cylinders.
NOTE
To thoroughly cover all surfaces of the cylinder interior, move the nozzle of the spray gun
from the top to the bottom of the cylinder. If
by accident the propeller is rotated following
this spraying, respray the cylinders to insure
an unbroken coverage on all surfaces.
g. Install lower spark plugs or install solid plugs,
and install dehydrator plugs in 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 exhaust 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 by inserting a protex plug
in the breather hose and clamping in place.
m. Seal all other engine openings exposed to atmosphere using suitable plugs or non-hygroscopic tape.
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 dehydrator 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 cylinder interiors
with corrosion-preventive mixture and replace all
desiccant and protex plugs.
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 o seal and cover
openings
d. Remove warning placards posted at throttle and
propeller
Remove and clean engine oil screen. then reinstall and safety. On aircraft that are equipped
with an external oil filter. install new filter element.
2-7
MODEL 210 & T210 SERIES SERVICE MANUAL
While these conditions are quite rare and will not
normally pose a problem to owners and operators,
they do exist in certain areas of the world and consequently must be dealt with when encountered.
f. Remove oil sump drain plug and drain sump.
Install and safety drain plug and fill engine with oil.
NOTE
Therefore, to alleviate the possibility of fuel icing
occurring under these unusual conditions it is permissible to add isopropyl alcohol or ethyelene glycol
monomethyl ether (EGME) compound to the fuel supply. See Figure 2-3 for fuel additive mixing ratio.
The corrosion-preventive mixture will mix
with the engine lubrication 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 dehydrator plugs and spark plugs or
plugs installed in spark plug holes and rotate
propeller by hand several revolutions to clear
corrosion-preventive mixture from cylinders.
i. Clean. gap and install spark plugs. Torque
plugs to value listed 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.
DELETED.
2-16.
SERVICING.
CAUTION
Diethylene glycol monomethyl ether
(DiEGME) has NOT been approved by engine
manufacturer for use with propeller single
engine aircraft
The introduction of alcohol or EGME compound into
the fuel provides two distinct effects: 1) it absorbs
the dissolved water from the gasoline and 2) alcohol
has a freezing temperature depressant effect.
Alcohol, if used, is to be blended with the fuel in a
concentration of 1% by volume. Concentrations
greater than 1% are not recommended since they can
be detrimental to fuel tank materials.
The manner in which the alcohol is added to the fuel
is significant because alcohol is most effective when
it is completely dissolved in the fuel. To insure
proper mixing the following is recommended.
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 BAYS. An area of each wing is sealed
to form an integral fuel bay. Recommended fuel
grades are listed in figure 2-4. Fuel bays should be
filled immediately after flight to lessen condensation
in bays and lines.
-~~~in bays and lines.
NOTE
Beginning with Serial 21064536, before refueling or when the aircraft is parked on a
slope, place the fuel selector handle in the
LEFT ON or RIGHT ON position, whichever
corresponds to the low wing. This will minimize crossfeeding from the fuller bay and .
reduce fuel seepage from the wing vents.
2-18A. USE OF FUEL ADDITIVES FOR COLD
WEATHER OPERATION. Strict adherence to recommended preflight draining instructions will eliminate
any free water accumulations from the tank sumps.
While small amounts of water may still remain in
solution in the gasoine, it will normally be consumed
and go unnoticed in the operation of the engine.
One exception co this can be encountered when operacing under the combined effect of: 1) use of certain
fuels, with 2) high humidity conditions on the ground
3; followed by flight at high altitude and low temperature. Under these unusual conditions small amounts
of water in solution can precipitate from the fuel
stream and freeze in sufficient quantities to induce
partial icing of the engine fuel system.
2-8
Revision 3
in
1. For best results the alcohol should be added
during the fueling operation by pouring the alcohol
directly on the fuel stream issuing from the fuel
nozzle.
2. An alternate method that may be used is to
premix the complete alcohol dosage with some fuel
a separate clean containerl dosage with some fuel
in a separate clean container (approximately 2-3
gallon capacity) and then transfer this mixture to the
^tank prior to the fuel operation.
Any high quality isopropyl alcohol may be used, such
as: Anti-icing fluid (MIL-F-5566) or Isopropyl alcohol (Federal Specification TT-I-735a).
Ethylene glycol monomethyl ether (EGME) compound
in compliance with MIL-1-27686 or Phillips PFA55MB, if used, must be carefully mixed with the fuel
in concentrations not to exceed 0.15o by volume.
ICAUTION1
Mixing of the EGiME compound with the fuel
is extremely important because concentration in excess of that recommended (0.15
percent by volume maximum) will result in
detrimental affects to the fuel tanks, such
as deterioration of protective primer and
sealants and damage to O-rings and seals
in the fuel system and engine components.
Use only blending equipment that is recommended by the manufacturer to obtain proper
proportioning.
Do not allow the concentrated EGIME compound to come in contact with the airplane
finish or fuel cell as damage can result.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
times.
When changing engine oil, remove and clean oil
pressure screen or install a new filter element on
aircraft equipped with an external oil filter. To
drain
as follows:
oil, proceed
a. Operate engine until oil temperature is at
normal operating temperature.
b. Remove oil drain plug from engine sump and
allow oil to drain into a container.
c. After engine oil has drained, install and safety
drain plug.
d. Remove and clean oil pressure screen or
change external oil filter element.
e.
Service engine with correct quantity and
viscosity of aviation grade engine oil.
NOTE
Refer to inspection charts for intervals
for changing engine oil and external filter
elements. Refer to figure 2-4 for correct
viscosities and capacities of aviation
grade engine oil.
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 over-stressed. 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. Under extremely dusty conditions, daily
servicing of the filter is recommended. To service
the induction filter, proceed as follows:
a. Remove
filter from
aircraft.
NOTE
Use care to prevent damage to filter
element when cleaning filter with
compressed air.
b. Clean filter by blowing with compressed air
(not over 100 psi) from direction opposite of normal
air flow. Arrows on filter case indicate direction of
normal air flow.
CAUTION
Do not use solvent or cleaning fluids to
wash filter.
Use only a water and
household detergent solution when
washing the filter.
c. After cleaning as outlined in step "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 assembly may be cleaned with
compressed air a maximum of 30 times or
it may be washed a maximum of 20
2-10
Revision 2
A new filter should be installed
after using 500 hours of engine operating
time or oneyear, whichever should occur
first.
However,
newexisting
filter should
installed
anytimea the
filter be
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 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.
f. 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.
2-22. VACUUM SYSTEM AIR FILTER. The vacuum
system central air filter keeps dust and dirt from
entering the vacuum operated instruments. Inspect
filter every 200 hours for damage. Replace filter
when damaged, every 500 hours of operation or whenever it becomes sufficiently clogged to cause suction
gage readings to drop below 4. 6 in Hg. Do not operate the vacuum system with the filter removed or a
vacuum line disconnected as particles of dust or other
foreign matter may enter the system and damage the
vacuum-operated instruments.
Excessive smoking will cause premature
filter clogging.
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) and clean water to neutralize
electrolyte or corrosion. Follow with a thorough
flushing with clean water. Do not allow
bicarbonate of soda to enter battery. Brighten
cable and terminal connection with a wire brush,
then coat with petroleum jelly before connecting.
Check the battery every 50 hours (or at 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.
MODEL 210 & T210 SERIES SERVICE MANUAL
2-24. TIRES. Maintain tire pressure at the value
specified in Section 1. When checking pressure,
examine tire for wear, cuts, bruises and slippage.
NOTE
Recommended tire pressure should be maintained. Especially in cold weather, remember
that any drop in temperature of the air inside
a tire causes a corresponding drop in 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. Remove valve cap and release all air.
b. Remove valve housing assembly.
c. Compress strut completely (stops in contact
with outer barrel hub).
d. Oil leveL
1. Fluid used should comply with Specification
MIL-H-5606.
2. Fill strut to bottom of valve installation
hole.
3. Maintain oil level at bottom of valve installation hole.
e. Fully extend strut.
f. Replace valve housing assembly.
g. With strut fully extended and nose wheel clear
of ground, inflate strut to 90 PSI.
NOTE
The nose landing gear shock strut will
normally require only a minimum
amount of service. Maintain the strut
extension pressure as shown in figure 11. Lubricate landing gear as shown in
figure 2-5. Check 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 lintwith hydraulic fluid
free clothSaturated
(MIL-H-5606)
or kerosene.
All surfaces
should be wiped free of excessive
hydraulic fluid.
2-26. NOSE GEAR SHIMMY DAMPENER. The
shimmy dampener should be serviced at least
every 100 hours. The dampener must be filled
completely with hydraulic fluid, free of entrapped
air with the compensating piston bottomed in the
rod. Check that piston is completely bottomed as
follows:
a. Remove shimmy dampener from the aircraft.
b. While holding the shimmy dampener in a
vertical position with the filler plug pointed
upward, loosen the filler plug.
c. Allow the spring to bottom out the floating
piston inside the shimmy dampener rod.
d. When the fluid stops flowing, insert a length of
stiff wire through the air bleed hole in the setscrew
at the end of the piston rod until it touches the
floating piston. The depth should be 3-13/16
inches.
NOTE
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
d. Reinstall filler plug and safety.
Wash a 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.
2-27. HYDRAULIC BRAKE SYSTEMS. Check
brake master cylinders and refill with hydraulic
fluid as specified in the inspection charts. 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 system.
2-28. LANDING GEAR HYDRAULIC RETRACTION
SYSTEM. Draining, filling and bleeding of the landing gear hydraulic system can be accomplished by
the following method.
a. Place aircraft master switch in OFF position
and place aircraft on jacks as shown in figure 2-2.
Bleed pressure from system by moving landing gear
selector valve to gear UP position.
selector valve to gear UP position.
CAUTION
Do not turn master switch ON while hydraulic
system is open to atmosphere. The pump
will automatically start, causing hydraulic
fluid to spray from any open line.
b. Drain system by removing cap from elbow on.
right side of power pack (behind access cover) and
attaching a drain hose to the elbow. Place end of
hose in a container of at least one gallon capacity
and using emergency hand pump, pump fluid into container. When power pack reservoir is empty, replace cap.
c. Fill power pack reservoir with MIL-H-5606 hvdraulic fluid by inserting a funnel or filler hose in
dipstick opening on top of power pack body.
2-11
MODEL 210 & T210 SERIES SERVICE MANUAL
d. Bleed system by cycling landing gear through
several cycles. Refill power pack reservoir with
MIL-H-5606 hydraulic fluid and remove aircraft
from jacks.
2-29. HYDRAULIC FLUID SAMPLING AND CONTAMINATION CHECK. At the first 50 and first 100
hour inspection and thereafter at each 500 hour inspection or one year, whichever should occur first,
a sample of fluid should be taken and examined for
sediment and discoloration. This may be done as
follows:
a. Place aircraft master switch in OFF position
and replace aircraft on jacks as shown in figure 2-2.
Bleed pressure from system by moving landing gear
selector valve to gear UP position.
CAUTION
Do not turn master switch ON while hydraulic
system is open to atmosphere. The pump
will automatically start, causing hydraulic
fluid to spray from any open line.
b. Remove cap from elbow on right side of power
pack (behind access cover) and place a nonmetal container below opening.
c. Place landing gear selector valve in DOWN position and operate emergency hand pump to pump fluid
into container.
d.
If the drain fluid is clear and not appreciably
darker in color than new fluid, continue to use the
present fluid.
e. If the fluid color is doubtful, place a fluid sample in a nonmetallic container and insert a strip of
polished copper in the fluid.
f. Keep copper in the fluid for six hours at a temperature of 70*F or more. A slight darkening of the
copper is permissible, but there should be no pitting
or etching visible up to 20X magnification. If pitting
or etching is evident, drain fluid from power pack
reservoir. Fill power pack with MIL-H-5606 hydraulic fluid and bleed air from system.
2-30.
OXYGEN SYSTEM.
2-31.
FACE MASKS.
2-32.
CLEANING.
Refer to Section 15.
Refer to Section 15.
2-33. 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.
2-34. UPHOLSTERY AND INTERIOR. Cleaning
prolongs the life of upholstery fabrics and interior
trim. To clean the interior, proceed as follows:
a. Empty all the ashtrays.
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.
2-12
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.
2-35. 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.
2-36. 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
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.
2-37. ALUMINUM SURFACES. The aluminum surfaces require a minimum of care, but should never
be neglected. The aircraft maybe 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.
2-38. PAINTED SURFACES. The painted exterior
surfaces of your new Cessna have a durable, long
lasting finish. Approximately 10 days are required
for the paint to cure completely; in most cases. the
curing period will have been completed prior to delivery of the airplane. In the event that polishing or
buffing is required within the curing period, it is
recommended that the work be done by someone ex-
MODEL 210 & T210 SERIES SERVICE MANUAL
perienced in handling uncured paint. Any Cessna
Dealer can accomplish this work.
W
^
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 chamots.
Harsh or abrasive soaps or detergents which cause
corrosion or scratches should never be used. Remove
stubborn oil and grease with a cloth moistened with
Stoddard solvent.
To seal any minor surface chips or scratches and
protect against corrosion, the airplane should be
waxed regularly with a good automotive wax applied
in accordance with the manufacturer's instructions.
If the airplane Is operated in a seacoast or other salt
water environment, it must be washed and waxed
more frequently to assure adequate protection. Special care should be taken to seal around rivet heads
and skin laps, which are the areas most susceptible
to corrosion. A heavier coating of wax on the leading
edges of the wings, and tail and on the cowl nose cap
and propeller spinner will help reduce the abrasion
encountered in these areas. Reapplication of wax will
generally be necessary after cleaning with soap solutions or after chemical de-icing operations.
2-39. ENGINE AND ENGINE COMPARTMENT. An
engine and accessories wash down should be accomplished during each 100-hour inspection to remove
might conceal component defects during inspection.
Also, periodic cleaning can be very effective in preventive maintenance.
Precautions should he taken when working with cleaning agents such as wearing of rubber gloves, an apron
or coveralls and a face shield or goggles. Use the
least toxic of available cleaning agents that will satisfactorily accomplish the work. These cleaning agents
include: (1) Stoddard Solvent (Specification P-D-680
type D), (2) A water alkaline detergent cleaner (MILC-25769J) mixed, 1 part cleaner, 2 to 3 parts water
and 8 to 12 parts Stoddard solvent or (3) A solvent
base emulsion cleaner (MIL-C-4361B) mixed 1 part
cleaner and 3 parts Stoddard solvent.
CAUTION
Do not use gasoline or other highly flammable
substances for washdown.
Perform all cleaning operations in well ventilated
work areas and ensure that adequate firefighting
and safety equipment is available. Do not smoke
or expose a flame, within 100 feet of the cleaning
area. Compressed air, used for cleaning agent,
application or drying, should be regulated to the
lowest practical pressure. Use of a stiff bristle
brush rather than a steel brush is recommended
if cleaning agents do not remove excess grease and
grime during spraying.
A recommended procedure for cleaning an engine and
accessories is as follows:
CAUTION
Do not attempt to wash an engine which is still
hot or running. Allow the engine to cool before
cleaning
a Remove engine cowling in accordance with Paragraph 12-3.
b. Carefully cover the coupling area between the
vacuum pump and the engine drive shaft so that no
cleaning solvent can reach the coupling or seal.
c. Cover the open end of the vacuum discharge tube.
d. Cover the vacuum relief valve filter, if installed
in the engine compartment.
e. Use fresh water for wash down when the engine is
contaminated with salt or corrosive chemicals. A
cleaning agent such as described previously may then
be used to remove oil and grime.
-_
CAUTION
Care should be exercised to not direct cleaning
agents or water streams at openings on the
starter, magnetos, alternator, vacuum pump
or turbocharger relief valve.
f. Thoroughly rinse with clean warm water to remove all traces of cleaning agents.
CAUTION
Cleaning agents should never be left on engine
components for an extended period of time.
Failure to remove them may cause damage
to components, such as neoprene seals and
silicone fire sleeves, and could cause additional
tioal corrosion.
corrosion.
g. Completely dry engine and accessories using
clean, dry compressed air.
h. Remove the cover over the coupling area.
i. Remove the cover from the vacuum discharge
tube.
j. Remove the cover from the vacuum relief valve
filter, if installed.
k. If desired, engine cowling may be washed with
the same cleaning agents, then rinsed thoroughly and
wiped dry. After cleaning engine, relubricate all
control arms and moving parts as required.
L Reinstall engine cowling.
WARNINGFor maximum safety, check that the magneto
switches are OFF, the throttle is closed, the
mixture control is in the idle cut-off position,
and the airplane is secured before rotating the
propeller by hand. Do not stand within the arc
of the propeller blades while turning the propeller.
m. Before starting engine rotate the propeller by
hand no less than four complete revolutions.
Revision 2
2-13
MODEL 210 & T210 SERIES SERVICE MANUAL
2-40. 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.
2-41. WHEELS. The wheels should be washed
periodically and examined for corrosion, chipped
paint, and cracks or dents in the wheel halves or in
the flanges or hubs.- If defects are found remove
and repair in accordance with Section 5. Discard
cracked wheel halves, flanges or hubs and install
new parts.
2-42.
LUBRICATION.
or under seacoast conditions, clean and lubricate
wheel bearings at each 100-hour inspection.
2-47. WING FLAP ACTUATOR. Clean and
lubricate wing flap actuator jack screw each 100
hours as follows:
a. Expose jack screw by operating flaps to fulldown position.
b. Clean jack screw threads with solvent rag and
dry with compressed air.
It is not necessary to remove actuator
from aircraft to clean or lubricate threads.
2-43. GENERAL DESCRIPTION. Lubrication
requirements
With oil can, apply light coat of No. 10 weight,
requirements are
are outlined
outlined in
in figure
figure 2-5.
2-5. Before
Before
non-detergent
oil to threads of jack screw
adding lubricant to a fitting, wipe the fitting free of non-detergent oil to threds of jack screw
dirt. Lubricate until grease appears around part
2-48. RODEND BEARINGS. Periodic inspection
being
wipe excess
excess grease
grease from
from
being lubricated
lubricated and
and wipe
and lubrication is required to prevent corrosion of
parts. The following paragraphs supplement
the bearing in the rod end. At each 100-hour
figure 2-5 by adding details not shown in the figure.
inspection,
control rods at the
inspection, disconnect
disconnect the
the control rods at the
2-44. NOSE GEAR TORQUE LINKS. Lubricate
aileron and inspect each rod end for corrosion. If
torque links every 50 hours. When operating in
no corrosion is found, wipe the surface of the rod
dusty conditions, more frequent lubrication is
end balls with general purpose oil and rotate ball
recommended , more frequent
freely to distribute the oil over its entire surface
and connect the control rods to the aileron. If
2-45.
TACHOMETER
DRIVE
SHAFT.
Refer
to
corrosion
is detected during inspection, install new
2-45. TACHOMETER DRIVE SHAFT. Refer to
Section 16
rod ends
2-46. 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 takeoff and landings are made. extensive taxiing is
required or the aircraft is operated in dusty areas
2-14
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
HYDRAULIC FLUID:
SPEC. NO. MIL-H-5606
OXYGEN:
SPEC. NO. MIL-0-27210.
SPECIFIED AVIATION GRADE FUELS;
WARNING
ONLY AVIATION GRADE FUELS ARE APPROVED FOR USE.
NOTE
APPROVED FUEL GRADES
ENGINE MODEL
Continental IO-520-L & TSIO-520-R
100LL (blue)
1
100 (green) (formerly 100/130)
1
NOTE
1.
Compliance with Continental Aircraft Engine Service Bulletin M82-8 and all supplements
or revisions thereto, must be accomplished.
SPECIFIED AVIATION GRADE OIL:
AVERAGE AMBIENT TEMPERATURE (°F) / OIL GRADE
|
SAE30
30
*
.
40º
30º
20º
10º
00
|
SAE
SAE 15W-50
600
50º
70°
SAE5
80°
90º
SAE 253-60
SAE 20W-50
Aviation grade ashless dispersant oil, conforming to Continental Motors Specification
MHS-24. and all revisions or supplements thereto, must be used except as noted in
paragraph 2-20, herein. Refer to Continental Aircraft Engine Service Bulletin M82-8,
and any superseding bulletins, revisions or supplements thereto, for further recommendations.
Oil capacities for the aircraft are given in the following chart. To minimize loss of
oil through the breather, fill to specified oil level on dipstick for normal operation
(flight of less than three hours duration). For extended flight, fill to FULL mark on
dipstick. Do not operate with less than MINIMUM FOR FLIGHT quantities listed. If
an external oil filter is installed, one additional quart of oil is required when filter
is changed.
CAPACITY
(TOTAL)
10
CAPACITY (TOTAL
WITH FILTER)
11
Figure 2-4.
2-16
Revision 2
NORMAL
OPERATION
MINIMUM
FOR FLIGHT
8
7
Servicing (Sheet 2 of 4)
MODEL 210 & T210 SERIES SERVICE MANUAL
.
DAILY
1
FUEL BAYS:
Service after each flight.
Keep full to retard condensation.
6
FUEL BAY SUMP DRAINS:
19
FUEL STRAINER:
15
OIL DIPSTICK:
Check on preflight. Add oil as necessary.
filler cap is tight and oil filler is secure.
Refer to paragraph 2-18 for details.
Drain off any water and sediment before first flight of the day.
Drain off any water and sediment before first flight of the day.
Refer to paragraph 2-20 for details.
8
PITOT AND STATIC PORTS:
Check for obstructions before first flight of the day.
7
OXYGEN CYLINDERS:
Check for anticipated requirements before each flight.
17
NOSE GEAR SHOCK STRUT:
Check that
Refer to Section 15 for details.
Check on preflight. Check inner barrel showing below outer barrel to be 1.00-2.00 (approximately 1.20) inches after bouncing. Deviation from these dimensions is cause to check and
service strut per paragraph 2-25.
25 HOURS
16
ENGINE OIL SYSTEM: FIRST 25 HOURS
Drain engine oil and change external oil filter (if equipped).
dispersant oil.
21
HYDRAULIC POWER PACK
Refill engine with ashless
Check every 25 hours and after a gear ext ension which uses the hydraulic hand pump.
50 HOURS
4
INDUCTION AIR FILTER:
Clean filter per paragraph 2-21.
13
BATTERY:
16
ENGINE OIL SYSTEM:
18
SHIMMY DAMPENER:
Check fluid level and refill as required in accordance with paragraph 2-26.
10
TIRES:
Maintain correct tire inflation as listed in Section 1. Refer to paragraph 2-24 for details.
17
NOSE GEAR SHOCK STRUT:
2
Replace as required.
Check electrolyte level and clean battery compartment each 50 hours or each 30 days.
Change oil each 50 hours if engine is NOT equipped with external filter; if equipped with
external oil filter, change oil and filter each 100 hours or every 6 months, whichever
occurs first.
Keep strut filled and inflated to correct pressure.
Refer to paragraph 2-25 for details.
HYDRAULIC FLUID RESERVOIR:
At first 50 and first 100 hours, thereafter at each 500 hours or one year, whichever comes
first, a sample of hydraulic fluid should be examined for sediment and discoloration as
outlined in paragraph 2-29.
Figure 2-4.
Servicing (Sheet 3 of 4)
2-17
MODEL 210 & T210 SERIES SERVICE MANUAL
D-
100 HOURS
2
HYDRAULIC FLUID RESERVOIR:
At first 50 and first 100 hours, thereafter at each 500 hours or one year, whichever comes
first, a sample of hydraulic fluid should be examined for sediment and discoloration as
outlined in paragraph 2-29.
3
FUEL/AIR CONTROL UNIT SCREEN:
Remove and clean screen.
5
VACUUM RELIEF VALVE FILTER:
Replace each 100 hours.
16
ENGINE OIL SYSTEM:
Change oil and filter each 100 hours or every 6 months, whichever occurs first.
19
FUEL STRAINER:
Disassemble and clean strainer bowl and screen.
200 HOURS
-
11
6
9
12
VACUUM SYSTEM CENTRAL AIR FILTER:
Inspect filter element for damage. Refer to paragraph 2-22.
FUEL BAY SUMP DRAINS:
Drain off any water or sediment.
FUEL RESERVOIR DRAIN:
Open drain valve(s) and drain off water and sediment.
BRAKE MASTER CYLINDERS:
Check fluid level and fill as required with hydraulic fluid.
<
l 11
>
500 HOURS
VACUUM SYSTEM CENTRAL AIR FILTER:
Replace every 500 hours. Refer to paragraph 2-22.
2
HYDRAULIC FLUID RESERVOIR:
At first 50 and first 100 hours, thereafter at each 500 hours or one year, whichever comes
first, a sample of hydraulic fluid should be examined for sediment and discoloration as
outlined in paragraph 2-29.
4
INDUCTION AIR FILTER:
Replace every 500 hours or annually.
A
14
Refer to paragraph 2-21.
AS REQUIRED
GROUND SERVICE RECEPTACLE
Connect to 24-volt, D.C. negative-ground power unit for cold weather starting and
lengthy ground maintenance of the aircraft's 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. Refer to Section 17.
Figure 2-4.
2-18
Revision 2
Servicing (Sheet 4 of 4)
MODEL 210 & T210 SERIES SERVICE MANUAL
FREQUENCY (HOURS)
METHOD OF APPLICATION
HAND
GREASE
GUN
OIL
CAN
WHERE NO INTERVAL IS SPECIFIED,
LUBRICATE AS REQUIRED AND
WHEN ASSEMBLED OR INSTALLED.
SYRINGE
(FOR POWDERED
GRAPHITE)
NOTE
The military specifications listed below are not mandatory,
but are intended as guides in choosing satisfactory materials.
Products of most reputable manufacturers meet or exceed
these specifications.
LUBRICANTS
PG
GR
GH
GL
OG
SS-G-659 .............
MIL-G-81322A ..........
MIL-G-23827A .....
MIL-G-21164C ..........
MIL-L-7870A ..........
PL VV-P-236 .............
GT ..............
OL VV-L-800A ............
POWDERED GRAPHITE
GENERAL PURPOSE GREASE
AIRCRAFT AND INSTRUMENT GREASE
HIGH AND LOW TEMPERATURE GREASE
GENERAL PURPOSE OIL
PETROLATUM
NO. 10WT NON-DETERGENT OIL
LIGHT OIL
.
NEEDLE BEARINGS
DAMPENER
PIVOTS
ALSO REFER TO
PARAGRAPH 2-44
OG
TORQUE LINKS
.
>;
^
S
.G
" ^
NEEDLE BEARING
(STEERING COLLAR)
"REFERTO PARA-
if//
\MAN GEAR
NOSE GEAR
j^\»\~
/-^MAIN
NOSE WHEEL
BEARINGS
Figure 2-5.
Mi
^^y
\ 6^
WHEEL
BEARINGS
/
N
-REFER TO
PARAGRAPH 2-47
Lubrication (Sheet 1 of 4)
2-19
MODEL 210 & T210 SERIES SERVICE MANUAL
DO NOT OIL IF OPERATING IN
EXTREMELY DUSTY CONDITIONS.
ELECTRIC FLAP
DRIVE MECHANISM
AILERON BELLCRANKS
ALSO REFER TO
PARAGRAPH 2-48
SCREW JACK
THREADS
ROD ENDS
NEEDLE
NEEDLE BEARING
ROLLERS
BEARINGS
6R
FLAP BELLCRANKS
AND DRIVE PULLEYS
CONTROL COLUMN
THRUST BEARINGS
ROD ENDS
NEEDLE BEARINGS
NEEDLE BEARINGS
6R NEEDLE BEARING
RUDDER BARS AND PEDALS
PARKING BRAKE
HANDLE SHAFT
BEARING BLOCK
OG
HALVES
GEAR WARNING AND
FUEL PUMP SWITCH
OILITE BEARINGS
(RUDDER BAR ENDS)
OL
ALL LINKAGE
POINT PIVOTS
OG
ENGINE CONTROLS
Figure 2-5.
2-20
Revision 2
Lubrication (Sheet 2 of 4)
MODEL 210 & T210 SERIES SERVICE MANUAL
SPRAY BOTH SIDES OF SHADED AREAS WITH
ELECTROFILM LUBRI-BOND "A" WHICH IS
AVAILABLE IN AEROSOL SPRAY CANS, OR
AN EQUIVALENT LUBRICANT. TORQUE
ATTACHING BOLT TO 10-20 LB-IN.
._-
-
NOSE GEAR
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-S-8660 silicone compound or equivalent lubricant,
applied sparingly to friction points, every 1000 hours or oftener if binding occurs.
No lubrication is recommended for the rotary clutch.
Apply DOOR-EZE lubricant to latch bolt.
Figure 2-5.
2-22
Lubrication (Sheet 4 of 4)
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
II
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.
a. 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 II 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 II
above. In addition, it is recommended that between annual inspections, all items be inspected
at the intervals specified in the inspection charts.
2-23
MODEL 210 & T210 SERIES SERVICE MANUAL
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 Cessna Parts
Distribution (CPD 2) through Cessna Service Stations.
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).
2. Aircraft Registration Certificate (FAA Form 8050-3).
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.
3. Pilot's Operating Handbook.
To be made available upon request:
1. Aircraft Log Book and Engine Log Book.
2-24
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
(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 Pilot's Operating Handbook. 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:
2-25
MODEL 210 & T210 SERIES SERVICE MANUAL
SPECIAL INSPECTION ITEM
EACH 200 HOURS
IMPORTANT
EACH 100 HOURS
READ ALL INSPECTION REQUIRE
MENTS PARAGRAPHS PRIOR TO
USING THESE CHARTS.
EACH 50 HOURS
PROPELLER
1.
Spinner
.......................
2.
Spinner bulkhead
3.
Blades ..
...................................
4.
Bolts and nuts
..
5.
Hub
6.
Governor and control
7.
Anti-Ice electrical wiring
8.
Anti-Ice brushes, slip ring and boots
.
..........................
. . ...
...
. ..
..
. . . ..
. . . .....
.
. . . . . . . . . . . .
........
..
..
...........
...........
.........
...............
..........
.......
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 crank shaft seal ...........
9.
Hoses, metal lines and fittings ................................
.
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
...................
..
.....
..
......
................
..
. . .. . . . . . . . . . . . . . . . . . . . . . . .
............
....
..............
..
................
3
10.
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
2-26
Revision 1
. .
.
..
..
..
. ..
. ..
.
.....
...........
. . . .....
•
..
. . . . ..
. ..
..
. ..
..
.
..
...................
. . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . ..
.
. . . . . . . . ..
. . . . . . ..
. ..
..
4
MODEL 210 & T210 SERIES SERVICE MANUAL
SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
6
18. Engine controls and linkage .....................................................
19. Engine shock mounts, mount structure and ground straps ...........................
20. Cabin heat valves, doors and controls
...........................................
21. Starter, solenoid and electrical connections ........................................
22. Starter brushes, brush leads and commutator ......................................
21
7
23. Alternator and electrical connections ..............................................
24. Alternator brushes, brush leads, 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 .........
..........................................
32. Engine cowling ................................................................
.
9
33. Turbocharger .................................................................
22
34. All oil lines to turbocharger waste gate and controller ................................
35. Waste gate, actuator and controller ...............................................
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 ...........................................
31
34
39. Fuel manifold valves, valve covers, and fuel system.................................
40. Fuel injection nozzles...........................................................
FUEL SYSTEM
23
27
1. Fuel strainer, drain valve and control, fuel bay vents, caps and placards ...............
2. Fuel strainer screen and bowl ....................................................
3. Fuel injector screen ............................................................
4. Fuel reservoir(s) ...............................................................
.
29
5. Drain fuel and check bay interior, attachment and outlet screens ......................
6. Fuel bays and sump drains .....................................................
D2057-3-13 Temporary Revision Number 8 - Apr 5/2004
© Cessna Aircraft Company
Revision 3
2-27
MODEL 210 & T210 SERIES SERVICE MANUAL
SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
7. Fuel selector valve and placards .................................................
8. Auxiliary fuel pump and throttle switches ..........................................
.....................
9. Engine-driven fuel pump ...............................
10. Fuel quantity indicators and sensing units .......................................
11. Fuel lines, check valve and vapor return line ...........
............................
24
.
12. Turbocharger vent system .......................................................
13. Engine primer .................................................................
14. Perform a fuel quantity indicating system operational test. Refer to
Section 16 for detailed accomplishment instructions ................................
·
32
LANDING GEAR
1. Brake fluid, lines and hose, linings, discs, brake assemblies and master cylinders .......
19
2. Main gear wheels ..............................................................
3. Wheel bearings................................................................
10
4. Main gear springs ..............................................................
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 ......................................
LANDING GEAR RETRACTION SYSTEM
NOTE
When performing an inspection of the landing gear retraction
system, the aircraft must be placed on jacks and an external power
source of at least 60 Amps should be used to prevent drain on the
aircraft battery when operating the system.
1. Operate the landing gear through five fault-free cycles ..............................
2. Check landing gear doors for positive clearance with any part of the
landing gear during operation, and for proper fit when closed. ........................
3. Check all hydraulic system components for security, hydraulic leaks and any apparent
damage to components or mounting structure .........
.......................
2-28
19
D2057-3-13 Temporary Revision Number 7 -.Oct 7/2002
© Cessna Aircraft Company
MODEL 210 & T210 SERIES SERVICE MANUAL
SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
Check doors, hinges, hinge pins and linkage for evidence of wear, other
.....................................
damage and security of attachment.
5. Inspect internal wheel well structure for cracks, dents, loose rivets, bolts
and nuts corrosion or other damage ..............................................
6. Check electrical wiring and switches for security of connections, and switch
operation. Check position indicator lights for proper operation.
......................
Check wiring for proper routing and support ..................
all systems and
of
rigging
proper
and
ensure
check
7. Perform operational
components including downlocks, uplocks, doors, switches, actuators
and power pack (observing cycle time). ...........................................
4.
8.
Check main gear strut to pivot attachment..........................................
Check condition of all springs. ...................................................
10. Hydraulic fluid contamination check .............................................
12
11. Clean power pack self-relieving check valve filter ..................................
12. Landing gear and door manifold solenoids (mounted on top of gear and
door manifolds) .............................................................
28
13. Hydraulic Pressure check primary and thermal relief valves and pressure switch. .......
30
9.
AIRFRAME
1. Aircraft exterior ................................................................
2. Aircraft structure ...............................................................
3. Windows, windshield, doors and seals ............................................
26
4. Seat stops, seat rails, upholstery, structure and mounting ............................
5. Seat belts and shoulder harnesses ...............................................
6. Control column bearings, sprockets, pulleys, cables, chains and turnbuckles ...........
7. Control lock, control wheel and control column mechanism ...........................
8.
Instruments and markings .......................................................
13
9. Vacuum system air filter.........................................................
10.
Magnetic compass compensation ................................................
11.
Instrument wiring and plumbing ..................................................
29
12. Instrument panel, shock mounts, ground straps, cover, decals and labeling.............
13. Defrosting, heating and ventilating systems and controls .............................
14. Cabin upholstery, trim, sun visors and ashtrays .....................................
15. Area beneath floor, lines, hose, wires and control cables .............................
16. Lights, switches, circuit breakers, fuses, and spare fuses ............................
Temporary Revision Number 7
7 October 2002
© 2002 Cessna Aircraft Company
Revision 3 2-29
MODEL 210 & T210 SERIES SERVICE MANUAL
SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
17. Exterior lights .....................................................................
18. 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 ............................................
.....................
23. Battery electrolyte ......................................
24. Emergency locator transmitter ...................................................
1.
14
15
25. Oxygen system ................................................................
26. Oxygen supply, masks and hose .................................................
16
27. De-ice system plumbing .........................................................
28. De-ice system components ......................................................
29. De-ice system boots ...........................................................
5
30. Vacuum Relief valve filter........................................................
33
31. Vacuum manifold check valve (If so equipped) ....................................
32. 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 .....................................
25
3. Trim control wheels, indicators, actuator and bungee ................................
4. Travel stops ...................................................................
5. Decals and labeling.............................................................
6. Flap control switch, rollers, tracks, and position indicator .............................
7. Flap motor, transmission, limit switches, structure, linkage, bellcranks etc ..............
8. Flap actuator jackscrew threads ............
................................
17
9. Elevator and trim tab hinges and push-pull tubes ...................................
10. Elevator trim tab actuator free play inspection ......................................
11. Elevator trim tab actuator lubrication inspection .....................................
12. Rudder pedal assemblies and linkage .........
18
18
.............................
13. External skins of control surfaces and tabs .........................................
2
14. Ailerons, hinges, and control rods .................................................
15. Internal structure of control surfaces ..............................................
16. Balance weight attachment ......................................................
2-30
Revision 3
2002 Cessna Aircraft Company
Temporary Revision Number 7
7 October 2002
MODEL 210 & T210 SERIES SERVICE MANUAL
SPECIAL INSPECTION ITEMS
1. First 25 hours: Use mineral oil confirming with MIL-C-6529 Type II for the first 25 hours 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 it is 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 for Special 100-hour inspection for 10-520
exhaust system. Refer to Section 12A for 50-hour inspection for turbocharged airplanes.
5. Change each 100 hours.
6. Each 50 hours for general condition and freedom of movement. These controls are not repairable.
Replace at each engine overhaul or sooner, if required.
7. Inspect each 50 hours.
8. Internal timing and magneto-to-engine timing limits are described in detail in Section 12.
9. Remove insulation blanket or heat shield and inspect for burned area, bulges or cracks. Remove
tailpipe and ducting; inspect turbine for coking, carbonization, oil deposits and impeller damage.
10. First 100 hours and each 500 hours thereafter. More often if operated under prevailing wet or dusty
conditions. Refer to Section 5 of this manual for inspection procedures.
11. If leakage is evident, refer to McCauley Governor Service Manual.
12. At first 50 hours, first 100 hours, and thereafter each 500 hours or one year, whichever occurs first.
13. Inspect for damage every 200 hours. Replace every 500 hours. Refer to paragraph 2-22.
14. Check electrolyte level and clean battery compartment each 50 hours or each 30 days, whichever occurs
first.
15. Refer to Section 17 of this manual.
16. Inspect masks, hose and fittings for condition, routing and support. Test, operate and check for leaks.
17. Refer to paragraph 2-47 for detailed instruction.
D22057-3-13 Temporary Revision Number 7 - Oct 7/2002
© Cessna Aircraft ComDanv
Revision 3
2-31
MODEL 210 & T210 SERIES SERVICE MANUAL
18. Replacement or overhaul of the actuator is required each 1,000 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.
19. Each 5 years, overhaul all retraction and brake system components. Check for wear, and replace all
rubber packings and backups and hydraulic hoses.
20. Refer to paragraph 2-48 for ball rod end inspection.
21. Refer to Section 17 of this manual for belt tension check procedures.
22. Replace check valve in the turbocharger oil line every 1,000 hours.
23. Beginning with T210, 21063661 and earlier airplanes modified by SK210-93. Check fuel strainer
insulation for security.
24. Beginning with T210, 21063661 and earlier airplanes modified by SK210-93. Check that the fuel line
insulation in the nose gear tunnel is in good condition. All fuel lines and vapor return lines are as far from
the exhaust system components as the installation will permit.
25. Compliance with Cessna Service Letter SE80-65 is required.
26. Inspect seat rails for cracks every 50 hours. Refer to Section 3.
27. Compliance with Cessna Single Engine Customer Care, Service Information Letter SE82-36 and Owner
Advisory SE82-36A is required.
28. Disassemble, clean and reassemble every 100 hours or 5 years, and whenever the solenoid is
accessible.
29. Each 1,000 hours, or to coincide with engine overhaul.
30. Can be operationally pressure checked in the airplane without power pack removal from the airplane
(refer to paragraph 5A-5A). To determine if the relief valve disassembly or adjustment is necessary,
relief valves can be bench checked after removal from power pack (refer to paragraph 5A-11A).
31. Each 100 hours or whenever fuel flow fluctuation is encountered, inspect fuel manifold valves, valve
covers, and fuel system components and lines for signs of leaks. Refer to Teledyne Continental Motors
Service Bulletin SB95-7.
32. Fuel quantity indicating system operational test is required every 12 months. Refer to Section 15 for
detailed accomplishment instructions.
33. Check condition and operation of check valve manifold, beginning five years from date of manufacture,
and every twelve months thereafter. Replace check valve manifold ten years from date of manufacture.
Refer to Airborne Products Reference Memo #39 for manufacture date information.
34. 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.
2-32
Revision 3
D2057-3-13 Temporary Revision Number 8 - Apr 5/2004
@Cessna Aircraft Company
MODEL 210 & T210 SERIES SERVICE MANUAL
2-45.
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 service life and time 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. 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.
2.
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
REPLACEMENT
TIME
COMPONENT
OVERHAUL
Restraint Assembly Pilot, Copilot,
and Passenger Seats
10 years
NO
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 Fluid
Carrying Teflon Hoses (CessnaInstalled) Except Drain Hoses
(Drain hoses are replaced
on condition)
10 years or engine overhaul,
whichever occurs first
(Note 1)
NO
Temporary Revision Number 7
7 October 2002
©2002 Cessna Aircraft Company
2-33
MODEL 210 & T210 SERIES SERVICE MANUAL
COMPONENT
Engine Compartment Flexible FluidCarrying Rubber Hoses (CessnaInstalled) Except Drain Hoses
(Drain hoses are replaced
on condition)
3.
REPLACEMENT
TIME
5 years or engine overhaul,
whichever occurs first
(Note 1)
OVERHAUL
NO
Engine Air Filter
500 hours or 36 months,
whichever occurs first
(Note 9)
NO
Engine Mixture, and Throttle,
Controls
At engine TBO
NO
Oxygen Bottle - Light Weight Steel
(ICC-3HT, DOT-3HT)
Every 24 years or 4,380 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 10)
NO
Standby Dry Vacuum Pump
500 hours or 10 years,
whichever occurs first
(Note 10)
NO
Check Valve (Turbocharger
Oil Line Check Valve)
Every 1,000 hours of
operation
(Note 11)
NO
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
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
12-34~~~~~~~~~~~
2-34
~~
@2002 Cessna Aircraft Company
OVERHAUL
Revision Number 7
~Temporary
7 October 2002
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
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 Corners 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 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.
Note 11: 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).
Temporary Revision Number 7
7 October 2002
© 2002 Cessna Aircraft Company
2-35
MODEL 210 & T210 SERIES SERVICE MANUAL
THIS PAGE INTENTIONALLY LEFT BLANK
2-36 2-36
© 2002 Cessna Aircraft Company
Temporary Revision Number 7
7 October 2002
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 3
FUSELAGE
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
FUSELAGE
............
.1C9/3-1
Windshield and Windows .....
. 1C9/3-1
Description ........
.1C9/3-1
Cleaning.
..........
1C9/3-1
Waxing
...........
1C9/3-1
Repair ...........
1C11/3-3
Scratches ......
. .1C11/3-3
Cracks .........
lC11/3-3
Sealing .........
. .1C14/3-6
Windshield. ..........
.. 1C14/3-6
Removal. ..........
1C14/3-6
Installation .....
..
.1C14/3-6
Windows
..........
. 1C14/3-6
Movable
...........
1C14/3-6
Removal and Installation
. 1C14/3-6
Wrap-Around Rear ......
1C14/3-6
Removal and Installation .. 1C14/3-6
Fixed ............
1C14/3-6
Cabin Structure ........
. 1C14/3-6
Sealing ...........
1C14/3-6
Cabin Doors .
.........
.1C14/3-6
Removal and Installation .- .. 1C14/3-6
|
Wedge Adjustment
..... 1C14/3-6
Weatherstrip
...
1C14/3-6
Sealing ...........
1C14/3-6
Latches .
..........
1C14/3-6
Description ...
. .1C14/3-6
Adjustment (Thru 21063640) 1C17/3-9
Lock ...........
1C17/3-9
Indexing Inside Handle (Thru
21063640) ....
. .1C17/3-9
Installation of Lock Assembly
(Beginning with 21063641) . 1C17/3-9
Installation of Latch Assembly
(Beginning with 21063641) .. 1C17/3-9
Installation of Cable Assembly
(Beginning with 21063641)
. 1C17/3-9
Rigging Cable Assembly
(Beginning with 21063641)
. 1C17/3-9
Rigging Inside Door Handle
(Beginning with 21063641) . 1C21/3-13
Door Pull Handle .
......
1C23/3-15
Removal and Installation . .1C23/3-15
Baggage Door.
........
C23/3-15
3-1.
FUSELAGE.
3-2.
WINDSHIELD AND WINDOWS..
(See figure 3-2.)
3-3. DESCRIPTION. The windshield and windows
are single-piece acrylic plastic panels set in sealing
strips and help in place by formed retaining strips
secured to the fuselage with screws and rivets.
Inmont Corp. 579.6 sealing compound used in con-
Removal and Installation .
. 1C23/3-15
Sealing ...
.
... 1C23/3-15
Scupper Drain Installation .
. 1C23/3-15
Seats .
..........
1C23/3-15
Pilot .
..... ...
. 1C23/3-15
Copilot ........
1C23/3-15
3rd and 4th ...
. .1C23/3-15
Description .......
1C23/3-15
Removal andInstallation. . 1C23/3-15
Bench (5th and 6th).
.....
1C24/3-16
Description .......
1C24/3-16
Removal and Installation. .1C24/3-16
Repair .
.........
1C24/3-16
Cabin Upholstery.
.........
1C24/3-16
Materials a d Tools ......
. 1D9/3-25
Soundproofing ...
........
1D9/3-25
Cabin Headliner ........
.. 1D9/3-25
Removal
..........
D9/3-25
Installation
.........
1D9/3-25
Upholstery Panels ........
D10/3-26
Removal and Installation . . . 1D10/3-26
Carpeting ...........
1D10/3-26
Removal and Installation
.. 1D10/3-26
Safety Provisions.
.......1D10/3-26
Baggage Retaining Net .....
1D10/3-26
Description .......
1D10/3-26
Safety Belts ........
.D10/3-26
Description . .....
.. D10/3-26
Shoulder Harness
......
1D11/3-27
Description .......
1D11/3-27
Inertia Reel Harness . . . D111/3-27
Description ......
1D11/3-27
Removal and Installation. 1D11/3-27
Glider Tow-Hook
.....
.. 1D11/3-27
Description ........
1D11/3-27
Rear View Mirror ..
. . . ..
D11/3-27
Description .........
1D11/3-27
. 1D11/3-27
Stretcher ........
Description ..
.....
1D11/3-27
Removal and Installation . . . 1D11/3-27
Cabin Step Installation ...
.
1.D13/3-29
Description .........
1D13/3-29
Removal and Installation . . . 1D13/3-29
Seat Rail Inspection ......
.
.1D14/3-30
junction with a felt seal, is applied to all edges of the
windshield and windows with exception of the wing
root area. The wing root fairing has a heavy felt
strip which 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
Revision 3
3-1
MODEL 210 & T210 SERIES SERVICE MANUAL
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. REPAIR. Replace extensively damaged transparent plastic rather than repair whenever possible,
since even a carefully patched part is not the equal
of a new section. either optically or structurally.
At the first sign of crack development, drill a small
end of the crack as shown in
hole at the extreme
figure 3-1. This serves to localize the cracks and
to prevent furtilzersplitting by distributing the strain
over a large area. If the cracks are small, stopping
them with drilledi holes will usually suffice until replacement or more permanent repair can be made.
The following repairs are permissible; however, they
are not to be located in the pilot's line of vision during landing or normal flight.
a. SURFACE PATCH. If a surface patch is to be
installed, trim away the damaged area and round all
corners. Cut a piece of plastic of sufficient size to
cover the damaged area and extend at least 3/4-inch
on each side of the crack or hole. Bevel the edges
as shown in figure 3-1. If the section to be repaired
is curved, shape the patch to the same contour by
heating it in an oil bath at a temperature of 248º to
302°F., or it may be heated on a hot plate until soft.
Boiling water should not be used for heating. Coat
the patch evenly with plastic solvent adhesive and
place immediately over the hole. Maintain a uniform
pressure of from 5 to 10 psi on the patch for a minimum of three hours. Allow the patch to dry 24 to 36
hours before sanding or polishing is attempted.
b. PLUG PATCH. In using inserted patches to
repair holes in plastic structures, trim the holes to
a perfect circle or oval and bevel the edges slightly.
Make the patch slightly thicker than the material
being repaired, and similarly bevel the edges. Install patches in accordance with procedure illustrated
in figure 3-1. Heat the plug until soft and press into
the hole without cement and allow to cool to make a
perfect fit. Remove the plug, coat the edges with
adhesive, and then reinsert in the hole. Maintain a
firm light pressure until the cement has set, then
sand or file the edges level with the surface; buff and
polish.
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 the scratch with a circular motion, keeping abrasive constantly wet with clean
water to prevent scratching the surface further. Use
minimum pressure and cover an area large enough
to prevent the formation of "bull's-eyes" or other
optical distortions.
.MIL-D-5549;
CAUTION
Do not use a coarse grade of abrasive.
320 is of maximum coarseness.
b.
finer grade abrasives until the scratches disappear.
c. When the scratches have been removed, wash
area thoroughly with clean water to remove all the
gritty particles. The entire sanded area will be
clouded with minute scratches which must be removed to restore the transparency.
d. Apply fresh tallow or buffing compound to a
motor-driven buffing wheel. Hold wheel against
plastic surface. moving it constantly over the damaged area until cloudy appearance disappears. A
2000-foot-per-minute 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 considerabley longer period
of time to attain the same result as produced by a buffing wheel.
e. When buffing is finished, wash the area thoroughly and dry with a soft flannel cloth. Allow surface to cool and inspect the area to determine if full
transparency has been restored. Apply a thin coat
of hard wax and polish the surface lightly with a clean
flannel cloth.
NOTE
Rubbing the plastic surface with a dry cloth
will build up an electrostatic charge which
attracts dirt particles and may eventaully
cause scratching of surface. After wax has
hardened, dissipate this charge by rubbing
the surface with a slightly damp chamois.
This will also remove dust particles which
have collected while the 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.
(See figure 3-1.)
a. When a crack appears in a panel, drill a hole at
the end of crack to prevent further spreading. The
hole should be approximately 1/8 inch in diameter,
depending on length of the crack and thickness of the
material.
b. Temporary repairs to flat surfaces can be accomplished by placing a thin strip of wood over each
side of the surface and inserting small bolts through
the wood and plastic. A cushion of sheet rubber or
aircraft fabric should be placed between the wood and
plastic on both sides.
c. A temporary repair can be made on a curved surface by placing fabric patches over the affected areas.
Secure the patches with aircraft dope, Specification No.
or lacquer, Specification No. MIL-L7178. Lacquer thinner, Specification No. Mil-T-6094
can also be used to secure the patch.
No.
Continue sanding operation, using progressively
3-3
MODEL 210 & T210 SERIES SERVICE MANUAL
15
19
NOTE
* When cabin top skin has been removed.
seal between skin (15) and radius formed
MODEL 210 & T210 SERIES SERVICE MANUAL
SEALING.
3-9.
3-10.
(See figure 3-2.)
WINDSHIELD.
(See figure 3-2.)
3-11. REMOVAL.
a. Drill out rivets securing top retainer strip.
b. Remove screws securing front retainer strip,
c. Remove wing fairings over windshield edges.
NOTE
Remove and tape compass and outside air
temperature gage clear of work area. Do
not disconnect electrical wiring.
3-18. FIXED. (See Figure 3-2.) Fixed windows.
mounted in sealing strips and sealing compound, are
held in place by various retainer strips. To replace
the side windows, remove upholstery and trim panels
as necessary and drill out the rivets securing retainers. Except for the left door, rear window and windshield, the aircraft is equipped with double windows.
Apply felt strip and sealing compound to all edges of
the window to prevent leaks. Check fit and carefully
file or grind away excess plastic. Use care not to
crack the window when installing.
3-19.
CABIN STRUCTURE.
3-20. SEALING.
d. Pull windshield straight forward, out of side
and top retainers.
3-12. INSTALLATION.
a. Apply felt strip and sealing compound or sealing
tape to all edges of windshield to prevent leaks.
b. Reverse steps in preceding paragraph for reinstallation.
c. When installing a new windshield, check fit and
carefully file or grind away excess plastic,
d. Use care not to crack windshield when installing.
Starting at upper corner and gradually working windshield into position is recommended.
3-13.
WINDOWS.
3-21.
(See figure 3-2.)
(See figure 3-2.)
CABIN DOORS.
(See figures 3-3 thru 3-4A.)
3-22. REMOVAL AND INSTALLATION. Removal of
cabin doors is accomplished by removing the screws
attaching the hinges and door stop, or by removing
the hinge pins attaching the door and door stop. If
permanent hinge pins are removed from the door
hinges, they may be replaced by clevis pins secured
with cotter pins, or new hinge pins may be installed
by inserting pin through both hinge halves and chucking a rivet set in a hand drill, hold one end of pin and
form head on opposite end. Reverse pin and repeat
process.
3-14. MOVABLE. (See figure 3-3.) A movable
3-14.window, hinged at the top, is installed in the left cabn
window, hinged at the top, is installed in the
.3-23. WEDGE ADJUSTMENT. Wedges, at upper forward edge of the door aid in preventing air leaks at
this point. They engage as the door is closed. Sev-
door on all aircraft and may also be installed in the
eral attaching holes are located in the wedges and the
right door as a customer option.
set of holes giving best results should be selected.
3-15. REMOVAL AND INSTALLATION.
a. Disconnect window stop (5).
b. Remove pins from. window hinges (6).
c. Reverse preceding steps for reinstallation. To
remove frame from plastic panel, drill out blind
rivets at frame splice. When replacing plastic panel
in frame, ensure sealing strip and an adequate coating of Presstite No. 579. 6 sealing compound is used
around all edges of panel.
3-16. WRAP-AROUND REAR. (See figure 3-2.)
The rear window is a one-piece acrylic plastic panel
set in sealing strips and held in place by retaining
strips.
3-17. REMOVAL AND INSTALLATION.
a. Remove upholstery as necessary to expose retainer strips inside cabin.
b. Drill out rivets as necessary to remove the retainers on both sides and the lower edge of window.
c. Remove window by starting at aft edge and pulling window into the cabin area.
d. Reverse preceding steps for reinstallation. Apply sealing strips and an adequate coating of sealing
compound to prevent leaks. When installing a new
window, check fit and carefully file or grind away excess plastic.
e. Use care not to crack the window when installing.
3-6
3-24. WEATHERSTRIP. Weatherstrip is bonded
around the edges of the cabin door and the movable
window opening. A hollow center, fluted type seal is
used. 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 seal at this point for drainage. Apply
a thin, even coat of EC-880 adhesive (3-M Co.) or
equivalent to each surface and allow to dry until tacky
before pressing into place.
3-25. SEALING.
(See figure 3-3.)
3-26. LATCHES. (See figure 3-4.)
3-26. LATCHES. (See figure 3-4.)
3-27. DESCRIPTION. (See figures 3-4, 3-4A and 3-5.)
Through 21063640, The cabin door latch is a pushpull bolt type, utilizing a rotary clutch for positive
bolt engagement. As the door is closed, teeth on underside of bolt engages 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. Beginning with 21063641, the rotary
clutch is replaced with a spring- loaded latch pin. As
the door is closed, (see figure 3-4A), push rod (14)
rides up on actuator (45), causing bolt (13) to disengage from catch (20), driving bolt into slot. As the
Door is opened, by pulling outboard on the handle (21),
bolt (13) is pulled out of slot, engaging spring-loaded
catch (20).
MODEL 210 & T210 SERIES SERVICE MANUAL
37
28
MODEL 210 & T210 SERIES SERVICE MANUAL
3-28.
ADJUSTMENT. (Thru 21063640.) (Refer to
figure 3-4.) Vertical adjustment of rotary clutch is
afforded by slotted holes which ensure sufficient gear-tobolt engagement and proper alignment. Adjustment for
bolt (2) extension is accomplished by loosening the four
bolt adjustment screws (26) sufficiently to move side bolt
guide (3) forward in the slotted holes to retract the bolt,
and aft to extend the bolt. Carefully close door after
adjustment to check bolt extension and clearance with
doorjamb and alignment with clutch assembly.
NOTE
Lubricate the door latch per Section 2. No
lubrication is recommended for the rotary
clutch.
3.29. LOCK. In addition to interior locks, a cylinder
and key type lock is installed on the left door. If the lock
is to be replaced, the new one may be modified to accept
the original key. This is desirable, as the same key is
used for the ignition switch and the cabin door lock.
After removing the old lock from door, proceed as follows:
a. Remove the lock cylinder from new housing.
b. Insert the original key into the new cylinder and file
off any protruding tumblers flush with cylinder. Without
removing key, check that cylinder rotates freely in the
housing.
c. Install the lock assembly in door, and check lock
operation with the door open.
d. Destroy the new key and disregard the code number
on cylinder.
3-30. INDEXING INSIDE HANDLE. (Thru
21063640.) (Refer to Figure 3-4.) When the inside
handle (12) is removed, reinstall in relation to position of
bolt (2), which should be in LOCK position, when
following these procedures.
a. Temporarily install inside handle (12) on shaft
assembly (16), aligning horizontally with arm rest.
b. Move inside handle (12) back and forth slightly to
ensure mechanism is centered in LOCK position.
c. Set inside handle adjustment screw (27) as required
to align handle parallel to centerline of handle axis.
d. Without rotating shaft assembly (16), remove
handle, and install placard (10) with LOCK index
forward and aligned horizontally with arm rest.
e. Install inside handle (12) to align with LOCK index
on placard (10), and install handle-retaining screw (13).
f. Ensure bolt (2) clears door post and teeth engage
clutch gear when handle is in CLOSE position.
3-31. INSTALLATION OF LOCK ASSEMBLY ON
LATCH ASSEMBLY. (Beginning with 21063641.)
(Refer to figure 3-4A.)
a. Assembly locking arm (3) with pin (5).
b. Place pin (5) in 1/8-inch hole of latch base assembly
(23).
c. Align .099-inch hole of locking arm (3) with .094-inch
| hole in latch base assembly (23), and install pin (4).
d. Assemble cam assembly (1) to locking arm (3). Cam
should be on latch side of locking arm (3).
e. Use washers between cam assembly (1) and cotter
pin (2), and install cotter pin on clevis bolt.
3-32. INSTALLATION OF LATCH ASSEMBLY.
(Beginning with 21063641.) (Refer to figure 3-4A.)
NOTE
Install with latch in CLOSED position.
a. Install latch assembly between door pan and door
skin.
b. Cable assembly should be forward of latch base
attach plate, and inboard of latch base cup.
c. Extend latch handle through cutout in door skin.
This will pull latch bolt back far enough to allow latch to
fall into place.
d. Push latch assembly aft so that bolt (13) and push
rod (14) extend through their respective holes.
e. Trip push rod (14) so that bolt (13) is fully extended
and outside handle (21) is flush.
f. Secure latch to door pan with four NAS220-5 screws
through base assembly (23) and two AN525-10R6 screws
through aft flange of door pan.
g. Drill eleven .128-inch holes to align with latch base
assembly (23).
NOTE
Do not oversize holes in the latch base, and do
not rivet base to skin at this time.
3-33. INSTALLING CABLE ASSEMBLY. (Beginning
with 21063641.) (Refer to figure 3-4A.)
NOTE
Remove cover assembly (41).
a. On pin end of cable assembly (25), attach clamp (26)
and self-locking clip-on nut (34), one-inch from end of
casing, as shown in Detail A.
b. Insert pin end of cable between door pan and door
skin at aft end of door. Push pin end of cable to top of
door.
c. Remove plug button (29) and align pin on cable with
pin guide (31), and insert pin through guide. Access is
gained through .875-inch hole (33).
d. Align clamp on cable casing with hole located oneinch below .875-inch hole (33), and install screw.
e. Check operation of cable. If sluggish operation of
cable is encountered, add S-1450-24A-0762 washers (27)
to self-locking clip-on nut (34) to facilitate smoother cable
operation.
NOTE
Washers are to be bonded to clip-on nut with
579.6 sealer (Inmont Corp., St. Louis, Missouri),
or equivalent.
3-34. RIGGING CABLE ASSEMBLY. (Beginning with
21063641.) (Refer to figure 3-4A.)
Revision 3
3-9
|
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Refer to paragraph 3-28
for bolt (2) adjustment.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
52
51
49
48
47
46
to correctly adjust striker plate (49), replace those (50)
shims with one #1212151-1 shim (53). When using one
#1212151-1 shim (53), also use two (50) shims, one between
the #1212151-1 shim (53) and the doorpost channel, and one
between the #1212151-1 shim (53) and striker plate (49).
Where two or more #1212151-1 shims (53) are needed, use
shims (50) as described above, plus, add one shim (50) between each #1212151-1 shim (53) used. In all cases, when
shimming striker plate (49), be sure to retain minimal distance between striker plate (49) and cabin door latch bolt.
Never grind the end of latch bolt to clear striker plate.
Always remove shims as required to maintain minimal
clearance.
NOTE
45.
46.
47.
48.
49.
50.
51.
52.
53.
Actuator
#1212150-1 Shim
Doorpost Jamb
Striker Plate Cover
Striker Plate
#1212147-1 Shim
Channel
Channel
#1212151-1 Shim
If cabin door is located forward such that the
door latch will not operate, this will not allow
the latch assembly push rod to ride up on the
actuator and trigger the latch bolt. Install
1212150 shims as required beneath the actuator, located on the cover assembly.
Figure 3-4A. Cabin Door and Latch Assembly (Sheet 3 of 3)
3-14
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Cabin door latch must be in OPEN position.
Latch must operate smoothly and freely.
1. Adjust striker plate (49) forward by installing 1212147-1 shims (50) as required, so that there
is a minimal clearance between bolt (13) and striker
(49).
NOTE
This adjustment will ensure that when the door
is opened from the outside, the bolt will engage
the latch catch, and the exterior handle will
stay open until the door is closed again.
NOTE
~~~NOTE ~3-37.
If cabin door is located too far forward such
that the door latch will not operate, this will
not allow latch assembly push rod (14) to
ride up on actuator (45) and trigger the latch
bolt (13), install 1212150-1 shims (46) as required beneath actuator (45), located on cover
assembly (48).
2. Close the cabin door from inside the aircraft.
When latch is overcentered, the exterior handle
should pull flush. If it does not pull flush, the connecting push rod from the door latch to the inside
handle assembly should be lengthened, adjus-
e. Rivet latch base (23) to door skin with MS20426A43 rivets.
f. Attach lock assembly casing (36) to door skin (37)
with nut (38) provided.
g. Install tumblers (35) and attach cam (1) to tumblers with screw and lock washer provided (40) and
(39)
h. Operate lock several times to assure that all
parts function properly.
NOTE
Steps "f", "g" and "h" apply to left-hand
doors only.
3-36. DOOR PULL HANDLE.
(See figure 3-3.)
REMOVAL AND INSTALLATION. (See figure 3-3, sheet 2.) The figure may be used as a
guide for removal and installation of the door pull
handle.
3-38. BAGGAGE DOOR (See figure 3-5.)
3-38. BAGGAGE DOOR
See
figure 3-5.)
3-39. REMOVAL AND INSTALLATION.
a. Disconnect door stop.
b. Remove hinge pin.
c. Reverse preceding steps for reinstallation.
3-40. SEALING.
(See figure 3-5.)
SCUPPER DRAIN INSTALLATION.
(See fig-
ted "out".
3-41.
3. On aircraft which have not been modified per
Mod Kit 1209062, when adjusting push rod (43), it
need only be adjusted 1/2 turn. To accomplish this,
base plate assembly (44) should be removed.
ure 3-5.)
a. Parts and materials required may be obtained
from the Cessna Supply Division.
b. Installation is accomplished with trim panel
under baggage door removed and carpet loosened
along left side of floor.
c. Remove sealant from intersection of bulkhead
(44), floor (45), and at lower left forward corner of
compartment for drain to lower fuselage.
d. Drill .250" drain hole (46) in lower left forward
corner of baggage compartment per detail F.
e. Install scupper (47) in lower left side of baggage
compartment by bonding scupper to floor and at both
ends with General Electric RTV-102 sealant.
f. Drill four number 40 holes through scupper (47)
and floor (45), equally spaced, starting 2.5" from
forward end. Install four sheetmetal screws (48).
g. Reinstall trim panel and carpet.
NOTE
When making this adjustment on the overcentering of the latch, it may be noticed
that there is a sharp, loud canning noise
when the inside handle is pushed down. It
is preferred that the outside door handle
be flush, even if the canning noise is
noticeable,
4. To make 1/2 turn adjustment, remove smaller
end of push rod (43) and turn it over (1800). Then reinstall base plate assembly.
5. When closing cabin door from the outside, by
using a large, sharp force on the outside handle, it is
possible to overcenter the inside handle, thus locking
one's self out To prevent this from occurring on
aircraft modified per Mod Kit 1209062, when adjust-
3-42. SEATS.
(See figure 3-6.)
3-43. PILOT. (See figure 3-6, sheet 1 of 3.)
a. Articulating recline/vertical adjust.
ing the push rod in step "2", adjust the push rod so
there is a sufficient force (6 to 12 pounds) against the
inside handle to prevent it from overcentering when
closing the door from the outside. (Refer to paragraph 3-35.)
6. Do not file, grind or sand any portion of the
bolt
7. Recheck clamps that secure cable. There
must not be any slippage between cable casing and
clamp.
8. After overcenter adjustment has been made,
install cotter pin (10) in clevis pin (9).
3-44. COPILOT. (See figure 3-6, sheet 1 of 3.)
a. Articulating recline.
b. Articulating recline/vertical adjust.
3-45. 3RD AND 4TH.
a. Articulating recline.
3-46. DESCRIPTION. These seats are manuallyoperated throughout their full range of operation.
Seat stops are provided to limit fore-and-aft travel.
3-47. REMOVAL AND INSTALLATION.
3-15
MODEL 210 & T210 SERIES SERVICE MANUAL
a. Remove seat stops.
b. Disengage the seat adjustment pin.
c. Slide seat fore-and-aft to disengage seat rollers
from rails.
d. Lift seat out.
e. Reverse preceding steps for reinstallation. Ensure all seat stops are reinstalled.
WARNING
It is extremely important that the pilot's seat
stops are installed. Acceleration and deceleration could possibly permit seat to become
disengaged from the seat rails and create a
hazardous situation, especially during take-off
and landing.
3-48. BENCH.
ure 3-6B.)
(See figure 3-6, sheet 3 of 3 and fig-
3-49. DESCRIPTION. These seats incorporate no
adjustment provisions and are bolted to the cabin
structure. The seat back folds down to provide
additional storage space on top of the main gear
wheel well and on top of the seat back. Beginning
with serial 21064773, the seat bottom may be removed from the frame by removing two bolts.
SHOP NOTES:
3-16
Revision 2
3-50. REMOVAL AND INSTALLATION.
a. Pull up on knob (1) to unlatch seat back.
b. Remove pin (10) from guide (8) on each side of
seat back.
c. Remove bolts (14) from the three seat legs.
d. Remove bolts (9) from both sides of seat bottom.
NOTE
Bolts (9) are located inside the main gear
wheel well.
e. With the seat back folded down, use care and
slide the two inside seat belts out from between the
seat back and bottom. Remove seat from aircraft.
f. Reverse preceding steps for reinstallation.
3-51. REPAIR. Replacement of defective parts is
recommended in repair of seats.
3-52. 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
MODEL 210 & T210 SERIES SERVICE MANUAL
1. Shim
22.
Baggage Door
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
Vertical Adjustment Handle
Fore/Aft adjustment Handle
Adjustment Pin
Spring
Seat Bottom
Articulating Adjustment Handle
Adjustment Screw
Bellcrank
Seat Back
Spacer
8
Channel
Torque Tube
Seat Structure
Roller
6
Stiffener
Trim
Seat Belt Retainer
Guide
Collar
* 21061574 THRU 21063178
21061574 THRU 21064135
*BEGINNING WITH 21064136
* 21061574 THRU 21062874
9
*2
4
19
2
Detail C
4
Detail A
11
Detail B
INFINITELY-ADJUSTABLE
PILOT AND COPILOT SEAT
Figure 3-6.
3-20
16
Seat Installation (Sheet 1 of 3)
12
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 4
WINGS AND EMPENNAGE
TABLE OF CONTENTS
WINGS AND EMPENNAGE. .......
Wings
............
Description ..........
Removal ............
Repair ...........
Installation ..........
Adjustment ..........
Vertical Fin ............
Description ..........
Removal
...........
Repair.............
Page No.
Aerofiche/Manual
.
1D20/4-1
1D20/4-1
1D20/4-1
D20/4-1
. 1 21/4-2
ID21/4-2
1 D21/4-2
1D21/4-2
. D21/4-2
1D21/4-2
1D21/4-2
4-1. WINGS AND EMPENNAGE.
4-2. WINGS.
Installation ........
..
D21/4-2
Horizontal Stabilizer ........
1D21/4-2
Description ..........
1D21/4-2
Removal ....
.......
1D22/4-3
Repair .
...........
1D22/4-3
Installation
.........
. D22/4-3
Stabilizer Abrasion Boots ......
1D22/4-3
Description
..........
1D22/4-3
Removal ...........
1D22/4-3
Installation ..........
1D22/4-3
(See figure 4-1.)
4-3. DESCRIPTION. Each all-metal wing panel is
a full cantilever type, with a single main spar, two
fuel spars, formed ribs and stringers. The front
fuel spar also serves as an auxiliary spar and provides the forward attachment point for the wing. An
inboard section of the wing, forward of the main spar,
is sealed to form an integral fuel bay area. Stressed skin is riveted to the spars, ribs and stringers
to complete the structure. An all-metal, balanced
aileron, flap, and a detachable wing tip are part of
each wing assembly. A navigation light is mounted
in each wing tip.
4-4. REMOVAL. Wing panel removal is most easily
accomplished if four men are available to handle the
wing. Otherwise, the wing should be supported with
a sling or maintanance stand when the fastenings are
loosened.
a. Remove wing gap fairings and fillets.
b. Drain fuel from wing being removed. (Observe
precautions outlined in Section 13.)
c. Remove cabin headliner in accordance with procedures outlined in Section 3.
WARNING
Oil, grease or other lubricants in contact
with high-pressure oxygen, create a serious fire hazard and such contact should be
avoided. Do not permit smoking or open
flame in or near aircraft while work is performed on oxygen systems,
d. (Refer to Section 15.) Rotate valves on three
cylinders clockwise to shut off filler line pressure;
the quick-release adapter on the cylinder-regulator
assembly will retain pressure within the cylinder.
Disconnect oxygen filler line at first tee upstream
from filler valve.
e.
Disconnect:
1. Electrical wires at wing root disconnects.
2. Fuel lines at wing root.
3. Pitot line (left wing only) at wing root.
4. Cabin ventilator hose at wing root.
5. Aileron carry-thru cable and aileron direct
cables of wing being removed, at turnbuckles behind
headliner front shield and doorpost shield.
NOTE
To ease rerouting the cables, a guide wire
may be attached to each cable before it is
pulled free from the wing. Then disconnect
cable from wire and leave the 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. If right wing is being removed, disconnect flap
cables from right flap drive pulley, and remove cable
guards and/or pulleys as required to pull flap cables
into right wing root area.
g. If left wing is being removed, relieve tension on
right flap cables at right flap drive pulley. Disconnect
right flap cables at flap actuator in left wing and remove pulleys to pull flap cables into left wing root
area.
NOTE
Rigging of flap actuator and components in
left wing need not be disturbed to remove
either wing. It is recommended that flap
be secured in streamlined position with
tape during wing removal to prevent damage,
since flap will swing freely.
h. Remove nut, washer and bolt attaching front fuel
spar to fuselage.
i. Remove bolts, washers and retainers holding
main spar dowel pins in position.
j. Support wing at inboard and outboard ends, and
4-1
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
remove dowel pins that attach main wing spar to
fuselage. It is recommended to remove the top dowel pin first, then lower outboard end of wing before
removing bottom dowel pin.
If a new wing is being installed, it will
be necessary to calibrate the fuel control
monitor in the cabin ceiling area. Refer
to Section 16 for calibration procedure.
NOTE
It may be necessary to use a long punch to
drive out main wing spar attaching dowel
pins, or to rock wing slightly while removing
pins. Care must be taken not to damage
dowel pins, spar fittings or spar carry-thru
fittings as these are reamed holes and close
tolerance dowel pins.
k.
NOTE
Be sure to install soundproofing panel in
wing gap before replacing fairing.
Remove wing and lay on padded stand.
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 by using the wing repair
jig, which may be obtained from Cessna. The 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.
h. Check operation of navigation, courtesy and
landing lights.
i. Check operation of fuel quantity indicator.
j. Install wing gap fairings and fillets.
k. Install headliner, interior panels, upholstery
and inspection plates.
1. Test operation of 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. Refer to Section 6 for
adjustment of aileron tabs.
INSTALLATION.
4-8.
VERTICAL FIN.
(See figure 4-2.)
NOTE
Refer to figure 4-1 for lubrication of dowel
pins prior to installation.
Hold wing in position with wing tip low.
Install:
1. Dowel pins attaching main spar to fuselage.
(Install bottom pin first, then rotate wing tip up, and
install top pin.)
2. Bolts, washers and nuts that hold main spar
attach dowel pins in position.
3. Front fuel spar attach bolt, washer and nut.
c. Route flap and aileron cables and make proper
connections.
d. Connect:
1. Electric wires at wing root disconnects.
2. Fuel lines at wing root.
3. Pitot line (if left wing is being installed. )
4. Cabin ventilator hose at wing root.
5. Oxygen filler line at tee in cabin top area.
a.
b.
-CAUTION
counterclockwise on
turn
valves
sure
to
Be
three oxygen cylinders to turn on filler line
pressure. Refer to Section 15 for a corplete oxygen system leak test prior to installing headliner.
e. Rig aileron system (Section 6).
f. Rig flap system (Section 7).
g. Refill wing fuel bays and check all connections
for leaks.
4-2
4-9. DESCRIPTION. The fin is primarily of metal
construction, consisting of ribs and spars covered
with skin. Fin tips are glass fiber/ABS construction.
Hinge brackets at the rear spar attach the rudder.
4-10. REMOVAL. The fin may be removedwithout
first removing the rudder. However, for access and
ease of handling, the rudder may be removed if desired, following the procedures outlined in Section 10.
a. Remove fairings on both sides 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 fin to fuselage.
d. Remove bolts attaching fin front and rear spars
to fuselage.
e. Remove fin.
4-11. REPAIR. A damaged fin may be repaired in
accordance with applicable instructions outlined in
Section 18.
4-12. INSTALLATION. Reverse procedures outlined in paragraph 4-10 to install the fin. Be sure to
check and reset rudder and elevator travel if any
stop bolts were removed or settings distrubed. Refer to Sections 8 and 10 respectively for setting elevator and rudder travel. Refer to figure 1-1 for
control surface travels.
4-13.
HORIZONTAL STABILIZER.
4-14.
DESCRIPTION.
(See figure 4-3.)
The horizontal stabilizer is
MODEL 210 & T210 SERIES SERVICE MANUAL
primarily of metal construction, consisting of ribs
and a front and rear spar which extends throughout
the full span of the stabilizer. The skin is riveted to
both spars and ribs. Stabilizer tips are constructed
of ABS. 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 an opening which
provides access to the elevator tab actuator screw.
Hinges on the rear spar support the elevator.
4-15. REMOVAL.
a. Remove elevators and rudder in accordance with
procedures outlined in Sections 8 and 10.
b. Remove vertical fin in accordance with procedures outlined in paragraph 4-10.
c. Disconnect elevator trim control cables at clevis,
turnbuckle and clamps inside tailcone, remove pulleys
which route the aft cables into horizontal stabilizer,
and pull cables out of tailcone.
d. Remove bolts securing horizontal stabilizer to
fuselage.
e. Remove horizontal stabilizer.
4-16. REPAIR. A damaged horizontal stabilizer
may be repaired in accordance with applicable instructions outlined in Section 18.
4-17. INSTALLATION. Reverse the procedures
outlined in paragraph 4-15 to install the horizontal
stabilizer. Rig the control systems as necessary,
following instructions outlined in applicable sections.
Set control surface travels to values listed in figure
1-1.
4-18.
STABILIZER ABRASION BOOTS.
NOTE
Accessory Kit AK182-217 is no longer available
from Cessna for installation of abrasion boots.
Order two abrasion boots (P/N 1232040-5) and
one cement (P/N EC1300LP), available from
Cessna Parts Distribution (CPD 2) through
immediately
Cessna Service Stations, for installation of
abrasion boots on aircraft not so equipped.
4-19. DESCRIPTION. The aircraft may be equipped
with two extruded rubber abrasion boots, one on the
leading edge of each horizontal stabilizer. These
edges
boots are installed to protect the stabilizer leading
edge from damage caused by rocks thrown back by
the propeller.
4-20. REMOVAL. The abrasion boots can be removed by loosening one end of the boot and pulling it
off the stabilizer with an even pressure. Excess adhesive or rubber can be removed with Methyl-EthylKeytone.
4-21. INSTALLATION. Install abrasion boots as
outlined in the following procedures.
a. Trim boots to desired length.
b. Mask off boot area on leading edge of stabilizer
with one-inch masking tape. allowing 1/4-inch margin.
c. Clean metal surfaces of stabilizer. where boot
is to be installed, with Methyl-Ethyl-Ketone.
d. Clean inside of abrasion boot with Methyl-EthylKetone and a Scotch Brite pad to ensure complete
removal of paraffin/talc. Then a normal wipe down
with MEK on a cloth will leave surface suitable for
bonding to the aluminum.
NOTE
Boots may be applied over epoxy primer,
but if the surface has been painted. the
paint shall be removed from the bond area.
This shall be done by wiping the surfaces
with a clean, lint-free rag, soaked with
solvent, and then wiping the surfaces dry,
before the solvent has time to evaporate,
with a clean, dry lint-free rag.
e. Stir cement (EC-1300, Minnesota Mining and
Manufacturing Co.) thoroughly.
f. Apply one even brush coat to the metal and the
inner surface of the boot. Allow cement to air-dry
fora minimum of 30 minutes, and then apply a second coat to each surface. Allow at least 30 minutes
(preferably one hour) for drying.
g. After the cement has thoroughly dried, reactivate the surface of the cement on the stabilizer, and
boot, using a clean, lint-free cloth, heavily moistened
with Toluol. Avoid excess rubbing, which would remove the cement from the surfaces.
h. Position the boot against leading edge, exercising
care not to trap air between boot and stabilizer.
NOTE
with a quick motion, and
with a quick motion, and
reposition it properly.
i. Press roll entire surface of boot to assure positive contact between the two surfaces.
. Apply a coat of GACO N700A sealer, or equivalent, conforming to MIL-C-21067 alon the tailing
of the bootto the surface of the skin to form a
neat, straight fillet
k. Remove masking tape and clean stabilizer of excess material
1. Mask to the edge of the boot for painting stabilize
Revision 3
4-3
MODEL 210 & T210 SERIES SERVICE MANUAL
4
2
3
-
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 5
LANDING GEAR. BRAKES AND HYDRAULIC SYSTEM (THRU 1978 MODELS)
IWARNING
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the propeller,
since a loose or broken wire or a component
malfunction could cause the propeller to rotate.
NOTE
Beginning with 1979 models, major changes were
made in the aircraft hydraulic system. To avoid
the confusion of serialization, Section 5A has been
added following this section. Section 5A covers
1979 and ON changes. However Section 5 contains
information which is still applicable to the aircraft
described in Section 5A. To avoid repetition of
information in Section 5A, the reader is referred
back to this section.
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
LANDING GEAR SYSTEM .........
1E7/5-3
Description .................
1E7/5-3
System Operation (Thru
21062273) .................
1E8/5-4
System Operation (Beginning
with 21062274) .............
1E8/5-4
Trouble Shooting ............
1E9/5-4
Main Landing Gear .............
1E16/5-12
Description
...............
1E16/5-12
Main Gear Strut Removal ....
17/5-13
Main Gear Strut Installation ..
1E17/5-13
Main Landing Gear Actuator .1E17/5-13
Removal .................
1E17/5-13
Disassembly ..............
1E17/5-13
Inspection of Parts ........
1E17/5-13
Parts Repair/Replacement .
1E19/5-15
Reassembly ..............
1E19/5-15
Installation ...............
1E20/5-16
Strut-to-Actuator Linkage ....
1E21/5-17
Description ...............
1E21/5-17
Pivot Assembly Removal ...
1E21/5-17
Pivot Assembly Installation
1E21/5-17
Main Gear Uplock Mechanism
1E21/5-17
Description ...............
1E21/5-17
Removal/Installation ......
1E21/5-17
Uplock Actuator ..........
1E21/5-17
Disassembly ...........
1E21/5-17
Inspection of Parts ......
1E22/5-18
Reassembly ............
1E22/5-18
Main Gear Downlock
Mechanism ................
1E23/5-19
Description ...............
1E23/5-19
Removal/Installation
of Components ...........
1E23/5-19
Downlock Actuator ........
1F1/5-20A
Disassembly ..........
1F1/5-20A
Inspection/Repair ......
1F1/5-20A
Reassembly ............
1F1/5-20A
Main Landing Gear Door System 1F2/5-21
Description
...............
1F2/5-21
Removal/Installation of
Strut and Wheel Doors ....
1F2/5-21
Strut Door Actuator
Removal/Installation .....
1F2/5-21
Strut Door Actuator
Disassembly ...........
1F2/5-21
Inspection ..............
1F2/5-21
Reassembly ............
1F2/5-21
Wheel Door Actuator
Removal ................
1F2/5-21
Disassembly (Thru
21062273) ............
1F2/5-21
Inspection (Thru
21062273) ............
1F2/5-21
Reassembly (Thru
21062273) ...........
1F4/5-22A
Disassembly (Beginning
with 21062274) ........
1F4/5-22A
Inspection (Beginning
with 21062274 ........
1F4/5-22A
Reassembly (Beginning
with 21062274) ........
1F4/5-22A
Main Wheel and Tire Assembly
1F4/5-22A
Description ...............1F4/5-22A
Removal ..................
1F4/5-22A
Cleveland Main Wheel
and Tire Disassembly ..
1F7/5-24
Inspection/Repair ....
1F7/5-24
Reassembly .........
1F9/5-26
McCauley Two-Piece
Main Wheel and Tire
Disassembly ..........
1F9/5-26
Inspection/Repair ....
1F9/5-26
Reassembly .........
1F9/5-26
McCauley Three-Piece
Main Wheel and Tire
Disassembly ..........
1F11/5-26B
Revision 3
5-1
MODEL 210 &T210 SERIES SERVICE MANUAL
Page No.
Aerofiche/Manual
TABLE OF CONTENTS
Inspection/Repair ....
Reassembly .........
Installation .........
Main Wheel Door Close System
Accumulator ...............
Description ...............
Removal
. ............
Disassembly/Reassembly ..
Installation ...............
Main Wheel Door Close System
Accumulator ...............
Description ...............
Removal ..................
Disassembly/Reassembly ..
Installation ...............
Main Wheel and Axle Removal
Installation ...............
Main Wheel Alignment .......
Wheel Balancing .............
Brake System ..................
Description ..................
Trouble Shooting ............
Brake Master Cylinder .......
Description ...............
Removal ..................
Disassembly ..............
Inspection/Repair .........
Reassembly ...............
Installation ...............
Hydraulic Brake Lines .......
Description ...............
Wheel Brake Assemblies .....
Description ...............
Removal ..................
Disassembly ..............
Inspection/Repair
.........
Reassembly ...............
Installation ...............
Brake Linings ............
Non-Asbestos Organic or
Metallic Brake Linings
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F12/5-27
1F14/5-29
1F14/5-29
1F14/5-29
1F14/5-29
1F14/5-29
1F16/5-31
1F16/5-31
1F16/5-31
1F17/5-32
1F17/5-32
1F17/5-32
1F17/5-32
1F17/5-32
1F17/5.32
1F17/5-32
1F17/5-32
1F17/5-32
1F17/5-32
1G8/5-46
1G8/5-46
1G8/5-46
1G8/5-46
1G8/5-46
1G8/5-46
1G10/5-48
1G10/5-48
1G10/5-48
1G10/5-48
1G10/5-48
1G10/5-48
1G10/5-48
Description
Operation ................
Removal ..................
Cleveland
Disassembly .........
1G11/5-49
1G11/5-49
1F19/5-34
Inspection/Repair
1G12/5-50
Brake System Bleeding ....
Parking Brake System .......
1F20/5-34A
1F20/5-34A
Description ...............
Removal/Installation of
1F20/5-34A
Components .............
1F20/5-34A
Revision 3
1G8/5-46
1F19/5-34
1F19/5-34
1F19/5-34
1F20/5-34A
1F20/5-34A
Inspection/Repair .........
Reassembly ...............
1G3/5-41
1G3/5-41
1G4/5-42
1G4/5-42
1G4/5-42
1G5/5-43
1G5/5-43
1G5/5-43
1G5/5-43
1G5/5-43
1G5/5-43
1G5/5-43
1G5/5-43
1G5/5-43
1G7/5-45
1G7/5-45
1G7/5-45
1G8/5-46
1G8/5-46
1F17/5-32
Checking Lining Wear .....
Brake Lining Installation ..
Inspection/Repair of
Components .............
Nose Gear System ..............
Description .................
Operation ...................
Trouble Shooting ............
Nose Gear Assembly Removal .
Disassembly of Nose Gear Strut
Inspection/Repair of Shock Strut
Reassembly .................
Installation .................
Shimmy Dampener ..........
Description ...............
Removal ..................
Disassembly ..............
5-2
1F11/5-26B
1F11/5-26B
1F11/5-26B
Torque Links ................
Description ...............
Removal ..................
Disassembly/Reassembly ...
Installation ...............
Nose Gear Uplock Mechanism .
Description ...............
Removal ..................
Installation ...............
Nose Gear Downlock Mechanism
Description ...............
Removal/Installation ......
Nose Gear Actuator ..........
Description ...............
Removal ..................
Disassembly ..............
Inspection/Repair of Parts ..
Assembly .................
Installation ...............
Uplock and Release Actuator
Removal/Installation ........
Disassembly/Inspection/
Repair/Reassembly .......
Nose Gear Door System .......
Description ...............
Operation ................
Removal/Installation ......
Door Mechanism Removal/
Installation ..............
Nose Gear Strut Door
Removal/Installation .....
Nose Wheel Steering System ..
Description ...............
Removal/Installation of
Components .............
Rigging ..................
Trouble Shooting ..........
Nose Wheel and Tire Assembly
1F22/5-36
1F22/5-36
1F22/5-36
1F22/5-36
1F22/5-36
1F22/5-36
1F24/5-38
1F24/5-38
1F24/5-38
1G2/5-40
1G2/5-40
1G2/5-40
1G2/5-40
1G2/5-40
1G3/5-41
1G3/5-41
...............
Reassembly
McCauley
1G10/5-48
....
.........
Disassembly .........
Inspection/Repair ....
Reassembly .........
Installation ...............
Hydraulic Power System
........
General Description ..........
Components Repair ..........
Repair Versus Replacement ...
Repair Parts and Equipment ..
Equipment and Tools .........
Hand Tools ...............
Compressed Air ...........
Hydraulic System Leak Check .
Power Pack ..................
Description ...............
On-Aircraft Hydraulic Power
Pack Operational Checks .
Removal ..................
Disassembly ..............
Inspection/Repair of
Components .............
Reassembly ...............
1G12/5-50
1G12/5-50
1G12/5-50
1G12/5-50
1G12/5-50
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G13/5-51
1G14/5-52
1G14/5-52
1G14/5-52
1G14/5-52
1G20/5-56
1G20/5-56
1G20/5-56
MODEL 210 &T210 SERIES SERVICE MANUAL
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
Installation of Power Pack
1G22/5-58
Pressure Switch ........... 1G22/5-58
Description ............
1G22/5-58
Disassembly ........... 1G22/5-58
Cleaning/Inspection/
Repair ...............
1G22/5-58
Assembly .............. 1G22/5-58
Adjustment ............
1G23/5-59
Relief Valve and Thermal
Relief Valve Assemblies .. 1G23/5-59
Bench Check of Relief Valve
1G23/5-59
and Thermal Relief Valve
Disassembly ...........
1G24/5-60
Inspection .............
124/5-60
Assembly and Adjustment 1G24/5-60
Door System Thermal
Relief Valve .............
1G24/5-60
Landing Gear and Door
Manifold Assemblies .....
1G24/5-60
Description ............
1G24/5-60
Solenoids ..............
1G24/5-60
Disassembly ........
1G24/5-60
Inspection/Cleaning of
1G24/5-60
Components .......
Assembly ........... 1G24/5-60
Landing Gear Manifold
(Thru 21062273) ......
1G24/5-60
1G24/5-60
Disassembly ........
1H1/5-60A
Inspection/Repair ....
Reassembly ......... 1H3/5-61
Landing Gear Manifold
(Beginning with 21062274) 1H3/5-61
1H3/5-61
Disassembly ........
Inspection/Repair ....
Reassembly .........
Adjustment ..........
Door Manifold Assembly
Disassembly .........
Cleaning/Inspection of
Components ........
Reassembly .........
Landing Gear Hand Pump
Description ..........
Removal ............
Disassembly .........
Inspection of
Components ........
Reassembly .........
Installation ..........
Landing Gear Position
Selector Valve ........
Removal/Installation .
Disassembly/Reassembly
Inspection of Parts ...
Strut Step Installation .....
Rigging Throttle-Operated
Microswitches ..........
Rigging of Main Landing Gear
Rigging of Nose Landing Gear
Rigging of Nose Gear Doors .
Rigging of Nose Gear Limit
Switches ................
Rigging of Nose Gear Squat
Switch ..................
Rigging Retractable Step
Cable Assembly ..........
Hydraulic and Electric
System Schematics .......
1H3/5-61
1H3/5-61
1H6/5-64
1H6/5-64
1H6/6-64
1H6/5-64
1H6/5-64
1H6/5-64
1H6/5-64
1H6/5-64
1H6/5-64
1H7/5-65
1H7/5-65
1H7/5-65
1H8/5-66
1H8/5-66
1H8/5-66
1H8/5-66
1H8/5-66
1H10/5-68
1H10/5-68
1H20/5-78
1H20/5-78
1H20/5.78
1H20/5-78
1H20/5-78
1H23/5-81
It is sometimes necessary to open the landing gear doors while the aircraft is on the ground with
the engine stopped. Operate the doors with the landing gear handle in the 'DOWN" position.
Except on aircraft 21062274 thru 21062954, to open the doors, turn off the master switch and
operate the hand pump until the doors are open. To close the doors, turn the master switch on.
On aircraft 21062274 thru 21062954, the hand pump is required to open and close the doors.
Position of the master switch for gear door operation is easily remembered by the following
rule: OPEN CIRCUIT = OPEN DOORS; CLOSED CIRCUIT = CLOSED DOORS.
WARNING
Before working landing gear wheel
wells, PULL-OFF hydraulic pump circuit
breakers. Thru Serial 21062273, the pump
circuit breaker is locaed in the circuit
breaker panel, located immediately forward
of the pilot's control wheel. Beginning with
Serial 21062274, the pump circuit breaker
is located in the circuit breaker panel,
located immediately forward of the left
forward doorpost. The hydro-electric power
pack system is designed to pressurize the
landing gear DOOR CLOSE sytem to 1500
PSI at any time the master switch is turned
on. Injury might occur to someone working
in wheel well area if master switch is
turned on for any reason.
5-1. LANDING GEAR SYSTEM.
5-2. DESCRIPTION. (Refer to Hydraulic and Electric
System Schematic, figure 5-37.) A hydraulically-
operated, retractable landing gear is employed on the
aircraft. The hydrdaulic power system includes
equipment required to provide a flow of pressurized
hydraulic fluid to the landing gear system. The
Cessna-manufactured, self-contained, hydro-electric
pack is located in the pedestal, with the hand pump
remotely located between the two front seats on the
floorboard. The gear selector handle is located on the
lower lefthand switch panel. A circuit breaker,
protecting the pump, is located in the circuit breaker
panel, located immediately forward of the pilot's
control wheel thru Serial 21062273. Beginning with
Serial 21062274, the pump circuit breaker is in the
circuit breaker panel, located immediately forward of
the left-hand forward doorpost. It is necessary to pull
out on the gear selector handle prior to moving the
handle up or down. The handle is fitted with a small
wheel for easy identification and assisting in holding
the handle in rough air. The right side of the pedestal
cover is fitted with a quick-removable access door for
checking and servicing the hydraulic fluid level. The
selector handle controls the gear position through an
electrical switch thru Serial 21062273 and by means of
a hydraulic shuttle valve on aircraft beginning with
Serial 21062274.
Revision 3
5-3
MODEL 210 & T210 SERIES SERVICE MANUAL
5-3.
SYSTEM OPERATION. ( Thru Serial 21062273 )
NOTE
Refer to the hydraulic schematic
diagrams at the end of this section to
trace the flow of hydraulic fluid as
outlined in the following paragraph.
5-3A. SYSTEM OPERATION. ( Beginning with Serial
21062274 )
NOTE
Refer to the hydraulic schematic diagrams
at the end of this section to trace the flow
of hydraulic fluid as outlined in the following paragraph.
When the aircraft master switch is closed, the
hydraulic power pack is ready to operate. When
the gear-up position is selected with the selector
switch, the gear valve solenoid connects the gearup line to the system pressure, and the gear-down
line to return. At the same time, the electric motor
that powers the hydraulic pump is turned on. The
hydraulic pressure is passed through a filter, and
is then divided between the gear valve and door
valve. Before hydraulic pressure can reach the
gear valve, a priority valve must open. The
priority valve can open only under two conditions:
1. There can be no pressure in the door close
line, because door close pressure is applied to a
piston to hold priority valve closed.
2. System pressure must build up to 750 psig
before the valve can open.
Pressure therefore, must go to the door-open line.
Pressure in the door-close line is prevented from
returning by the door-close lock check valve. and
the valve is opened by a piston that senses dooropen pressure. When the presure reaches 400 psig,
the door-close lock check valve opens and the
doors on the aircraft open. At 750 psig. the
priority valve opens and the landing gear begins to
retract. As soon as the landing gear is locked in
the UP position. the landing gear up limit switches
sequence the door solenoid valve to the door close
position. When pressure in the door-close line
reaches 1500 psig. the pressure switch shuts off the
motor and the GEAR-DOWN cycle is similar to the
cycle. except
except. the gear solenoid
solenoid is not
GEAR-UP cycle,
energized
the gear-down
during
cycle. The system
When the aircraft master switch is closed, the hydraulic power pack is ready to operate. When the
gear-up position is selected with the selector handle
the selector valve connects the gear-up line to the
system pressure, and the gear-down line to return.
At the same time, the electric motor that powers the
hydraulic pump is turned on. The hydraulic pressure
is passed through a filter, and is then divided between
the selector valve and door valve. Before hydraulic
pressure can reach the selector valve, a priority
valve must open. The priority valve can open only
under two conditions:
1. There can be no pressure in the door close
line, because door close pressure is applied to a
piston to hold priority valve closed.
2. System pressure must build up to 750 psig
before the valve can open. Pressure therefore, must
go to the door-open line. Pressure in the door-close
line is prevented from returning by the door-close
lock check valve, and the valve is opened by a piston
that senses door-open pressure. When the pressure
reaches 400 psig, the door-close lock check valve
opens and the doors on the aircraft open. At 750 psig,
the priority valve opens and the landing gear begins to
retract. As soon as the landing gear is locked in the
UP position, the landing gear up limit switches sequence the door solenoid valve to the door close position. When pressure in the door-close line reaches
1500 psig, the pressure switch shuts off the motor and
the GEAR-DOWN cycle is similar to the GEAR-UP
cycle. The system has been designed so that at any
conditions, thefirst operation of the system after
then move the gear into the newly-selected position,
has been designed so that at any time during
system operation. the direction of system of
operation may be reversed. Under these
the selector switch is moved is to completely open
the doors, and then move the gear into the newlyselected position. after which, the doors will close
again. There is no danger of interference between
the gear and doors of the aircraft, since the gear
does not receive hydraulic pressure unless the
doors are in the fully-opened position.
SHOP NOTES:
5-4
time during system operation, the direction of system
of operation may be reversed.
Under these conditions,
the first operation of the system after the selector
handle is moved is to completely open the doors, and
after which, the doors will close again.
There is no
danger of interference between the gear and doors of
the aircraft, since the gear does not receive hydraulic pressure unless the doors are in the fully-opened
position.
MODEL 210 & T210 SERIES SERVICE MANUAL
5-4.
TROUBLE SHOOTING.
Just because this chart lists a probable cause, proper checkout procedures cannot be deleted and the replacement of a part is not necessarily the proper solution to the problem. The mechanic should always look for obvious problems such as loose or broken parts, external leaks, broken wiring, etc. To find the exact cause of a
problem, a mechanic should use a hand pump, pressure gage and a voltmeter to isolate each item in the system.
Hydraulic fluid will foam if air is pumped into system, causing fluid to be blown overboard thru pack vent line.
The problems listed are all with the systems controls in their normal operating position: Master switch ON,
hydraulic pump breaker IN and landing gear breaker IN. During landing gear system servicing, a power
supply capable of maintaining 27. 5 volts throughout the gear cycle must be used to augment the ship's battery.
CAUTION
Prior to using Hydro-Test unit with power pack, remove and dry off
filler plug and dipstick. Adjust cap tension so that no movement
of cap is apparent. Failure to accomplish these procedures could
result in filler cap coming loose from power pack.
TROUBLE
MOTOR PUMP WILL NOT
OPERATE GEAR BUT
EMERGENCY HAND PUMP
WILL OPERATE GEAR.
PROBABLE CAUSE
REMEDY
Low voltage (in flight).
Check alternator and wiring.
Fluid level low in reservoir.
Refill reservoir.
Motor pump failure.
Replace pump.
Faulty check valve
Replace valve
_
Loose or clogged suction screen
assembly in power pack
Remove power pack, disassemble
and clean suction screen. Check
screen for contamination. determine cause of contamination and
remedy. Replace screen assembly or seal existing assembly.
Prime parts to be assembled
with Grade T Primer, using care
to avoid getting primer on screen.
Seal with hydraulic sealant ( Catalog #69; Loctite Corp.) upon
installation. Allow 15-30 minutes
cure time if primed; 2-4 hours
if unprimed.
NOTE
Motor and pump are not repairable and must be replaced.
Pump frozen.
Remove motor and coupling
from top of power pack and
replace pump.
Broken pump or motor drive
shaft or coupling.
Remove motor and pump from
top of power pack and replace
motor, pump and coupling.
5-5
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont.)
TROUBLE
MOTOR PUMP WILL NOT
OPERATE GEAR BUT
EMERGENCY HAND PUMP
WILL OPERATE GEAR (Cont).
PUMP OR EMERGENCY PUMP
WILL NOT BUILD PRESSURE_
IN SYSTEM.
DOORS WILL NOT CLOSE
GEAR INDICATOR LIGHT
NOT ILLUMINATED.
5-6
PROBABLE CAUSE
REMEDY
If motor was not turning,
check wiring and motor.
Check motor for loose or broken
connections; check for frozen
pump or coupling. Check
circuit breaker in pedestal.
Bad pump shaft seal.
Replace pump.
External leakage around top
of pump assembly
Remove motor and pump assemblies from top of power pack and
replace upper packing and/or
back-up rings
Air lock in pump (new pack
installation or pump replacement.
Remove filter and intermittenly
bump start switch until fluid flows.
Replace filter.
Bad pump body O-rings
Remove motor and pump assemblies from top of power pack and
replace lower packing and/or
back-up rings
No fluid in reservoir.
Refill reservoir.
_
Broken hydraulic line.
Check for evidence of leakage
and repair or replace line.
Flush out system and refill
reservoir.
Filter in outlet check valve improperly positioned in filter
body, or seal between filter
and check valve improperly
positioned.
Replace seal and position
filter in retainer with
Petrolatum.
Bad O-ring actuator
piston; O-ring left out
after repair.
Disconnect line upstream from
actuator and check for pressure.
Perform this check for all
actuators in system.
Bad O-ring on priority valve in
gear manifold assembly. 0ring left out or damaged during
repair of valve.
Disassemble manifold and
replace O-ring.
Bad O-ring on gear or door
control valve.
Replace O-ring.
Thermal relief valve stuck open.
Replace valve.
Master switch not on.
Turn master switch on.
Broken or loose door
close hydraulic line.
Locate and repair or
replace defective line.
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont)
TROUBLE
DOORS WILL NOT CLOSE
GEAR INDICATOR LIGHT
NOT ILLUMINATED. (Cont)
PROBABLE CAUSE
REMEDY
Defective limit switch circuit.
Check limit switch settings; locate
and repair or replace limit switch
circuit.
Landing gear did not lock
into position.
Check landing gear uplock
and/or downlock mechanism
for proper operation.
Broken ground wire at socket
Repair or replace wire; check
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
REMEDY
GEAR UNLOCKS BEFORE DOORS
ARE FULLY OPEN
Restriction in door open or door
close line.
Using pressure gage, check pressure in door open or door close line,
when gear unlocks. If pressure is
greater than 700 psi, check for restrictions. Locate restrictions and
remove. If contaminates are in line,
investigate cause and remedy; flush
system.
DOORS OPEN BUT GEAR
DOES NOT OPERATE.
Improper wiring.
Check circuitry, using wiring
diagrams in this Section
or Section 20.
Gear solenoid jammed or
stuck ( Thru Serial 21062273 )
Disassemble valve and
replace defective parts.
Shorted gear control switch.
( Thru Serial 21062273 )
Check switch circuitry.
Priority valve setting too high
or stuck closed.
Check valve componets for
defects. Replace as necessary.
Faulty O-rings downstream
of priority valve (anywhere
in system).
Locate faulty unit and replace
0-rings.
DOORS OPEN BUT GEAR
DOES NOT OPERATE (DOWN
AND LOCKED ONLY).
Faulty or stuck squat switch.
Check switch wiring or setting.
HAND PUMP DOES NOT BUILD
PRESSURE, BUT ELECTRIC
PUMP OPERATES PROPERLY.
Check valve in hand pump
sticking.
Inspect check valve.
Defective hand pump outlet check
valve.
Replace valve.
Main gear or downlock actuator
O-ring leaking.
Disassemble actuator and
replace O-rings.
Fluid level low in reservoir.
Refill reservoir.
Downlock rod adjustment
incorrect (mainly LH rod).
Adjust rod end to lengthen
actuator one turn.
Pump failure.
Replace pump.
Low voltage in electrical system.
Check alternator and wiring.
Pump motor brushes worn.
Replace pump motor.
Downlocks not in full unlock
position.
Adjust downlocks.
Fluid leak in door or gear line.
Locate and repair or replace
broken line or fitting
LANDING GEAR OPERATION
EXTREMELY SLOW.
5-8
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont)
TROUBLE
LANDING GEAR OPERATION
EXTREMELY SLOW (Cont)
PROBABLE CAUSE
REMEDY
Air leakage around pump
suction screen assembly.
Either replace suction screen
assembly or seal and install
existing assembly as follows:
Prime parts to be assembled
with Grade T Primer, using
care to avoid getting primer
on screen. Seal with hydraulic
sealant (Catalog #69; Loctite
Corp.) upon installation. Allow
15-30 minutes cure time if
primed; 2-4 hours if unprimed.
Defective piston seal in gear
or door cylinder.
Replace with new seal.
Excessive internal power pack
leakage.
Remove and repair or replace
power pack.
PUMP OPERATES, DOORS OPEN
AND GEAR STARTS TO EXTEND.
DOORS CLOSE BEFORE GEAR
IS COMPLETELY EXTENDED;
HAND PUMP WILL NOT PUMP
GEAR DOWN.
Downlock switch makes contact before gear is down and
locked.
Reset downlock actuator switches;
replace if damaged.
Interference between downlock
and gear saddle clamp
bolt head.
Remove interference.
POWER PACK EXTERNAL
LEAKAGE.
Static seals (all fittings).
Remove and replace O-rings
and/or back-up rings as
required. Check tubing
flares for leaks.
Gear or door solenoid.
Replace O-rings.
Transfer tubes between manifold
and power pack body.
Disassemble power pack and
replace O-rings.
Reservoir cover.
Remove power pack and remove
cover; replace seals.
GEAR DOWN-LOCK WILL NOT
RETURN TO FULL-LOCK
POSITION.
Binding in spring and
tube assemblies.
Check operation to locate
binding and eliminate.
DOORS CLOSE BEFORE ALL
GEARS ARE FULLY LOCKED.
Faulty limit switch.
Replace switch_
Short in wiring.
Check wiring continuity.
Cracked terminal block.
Replace terminal block.
Lines between downlock
actuators crossed.
Properly route lines.
Lines crossed at gear uplock
valve.
Properly route lines.
DOORS WILL OPEN BUT
GEAR WILL NOT RETRACT.
Gear uplock valve installed backward. Install properly.
5-9
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont).
TROUBLE
MALFUNCTION OF GEAR
INDICATOR LIGHTS.
PROBABLE CAUSE
1.
2.
Both lights on at same time.
Light will change from green
to amber or in reverse when
gear control switch is moved.
SYSTEM WORKS NORMALLY EX- Leak in door close system.
CEPT MOTOR TURNS ON AND
OFF AT REGULAR INTERVALS.
(GEAR IN EITHER UP OR DOWN
POSITION). GEAR DOORS SAG
WHILE AICRAFT IS ON
GROUND. ENGINE AND ELECTRICITY OFF.
REMEDY
Check ground wire for proper
connection.
Refer to the following procedure and to figures 5-27
and 5-33A.
1.
Support aircraft on jacks or secure tail in the event something might unlock nose wheel and allow
it to collapse.
2.
Remove console cover and sheet metal cover from power pack support.
3. Master switch OFF.
4. Remove cap from pressure port on pedestal structure and install pressure gage to port.
5. Open doors as required to bleed any pressure in system.
6. Remove hand pump line from power pack port fitting (left-hand aft fitting).
7.
Attach flex line to disconnected line.
(have port open)
8. Remove door close line from its fitting on power pack (left hand forward fitting).
9. Connect flex line to door close port (fitting) on power pack and pressurize to 1500 psi with hand pump.
10. Observe pressure gage for leak-down; pressure should hold for better than 10 minutes.
(a) Master switch OFF - if leakage comes from hand pump fitting (open) 3 or 4 drops thermal relief valve leaking; replace.
(b) No leaks above - pull hydraulic circuit breaker out, master switch ON - repressurize
system with hand pump to 1500 PSI.
1. If hand pump port leaks in this configuration, lock out valve is leaking.
11.
With the preceding checks completed, and whether leaks were found or not, make this final
check while working in this area:
Remove flex line from door fitting and attach to door line and apply pressure to system.
There might be a alight bleed-down on first application of pressure pump to 1500 PSI
a second time. Pressure should hold.
12. The preceding procedure checks the door cylinders for leakage. If the system does not bleed down,
disconnect added equipment and reconnect lines and pressure cap to pressure port and reinstall
console covers. If on this last test, pressure does not hold, one or more of the door cylinders
are leaking. They will have to be checked individually. TEST SYSTEM BEFORE FLIGHT.
Revision 2
5-11
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont).
TROUBLE
UNEVEN FALL OF MAIN GEAR.
PROBABLE CAUSE
Air in system.
Bleed system of air.
Cold operating temperatures.
Operate power pack untilfluid
has reached operating temperature.
Improper snubber adjustment.
Adjust flow control valve
in gear manifold.
5-5. MAIN LANDING GEAR.
5-6. DESCRIPTION. The tubular main landing
gear struts rotate aft and inboard to stow the main
SHOP NOTES:
5-12
REMEDY
wheels below the baggage compartment. Struts
are down locked by an overcenter lock, actuated by
a hydraulic cylinder for each strut. Uplocks are
located on the main wheel stowage bay forward
MODEL 210 & T210 SERIES SERVICE MANUAL
bulkhead. Uplocking the gear pawls here, hold the
struts in the stowed position. Rotation of the
landing gear to extend or retract the struts is
achieved through pivot assemblies, which are in
turn bolted through a splined shaft, to the
hydraulic rotary actuators.
5-7. MAIN GEAR STRUT REMOVAL. (See figure
5-1.)
a. Jack aircraft in accordance with procedures
outlined in Section 2.
b. Disconnect brake line (17) at wheel cylinder and
drain brake system of strut being removed,
c. Place landing gear handle up, with master switch off,
and operate emergency hand pump until main gear
downlocks release.
d. Remove bolt (31) and nut securing strut to pivot
assembly (3).
e. Work strut and wheel from pivot assembly (3).
5-8. MAIN GEAR STRUT INSTALLATION. (Refer to
figure 5-1.)
5-10. REMOVAL OF MAIN GEAR ACTUATOR.
a. Remove seats and peel back carpet as
necessary to gain access to plate above actuator:
remove access plate.
b. Remove access plate from bulkhead forward of
actuator.
c. Disconnect and drain hydraulic brake line at
wheel brake cylinder.
d. Place landing gear control handle UP. with
master switch off. and operate emergency hand
pump until main gear downlocks release.
e. Disconnect and cap or plug all the hydraulic
lines at the actuator.
f. Remove bolts attaching actuator mounting
flange to bulkhead forging.
g. Work actuator free of forging and pivot
assembly, remove actuator.
DISASSEMBLY OF ACTUATOR. (Refer to
5-11.
figure 5-2.)
NOTE
Leading particulars of the actuator are as
follows:
NOTE
The following procedure installs the landingder
gear as a complete assembly. Refer to
applicable paragraphs for installation of
individual components.
a. Lubricate new O-rings (19) ad end
d of strut (5) with
Petrolatum W-P-236, hydraulic fluid MIL-L-5606, or
Corning DC-7 (keep DC-7 away from areas to be painted)
before installation. Install O-rings (19) on plug (20).
b. Remove caps from brake line fitting (18) and brake
line (17), attach brake line (17) to brake line fitting (18),
and work plug (20) and strut (5) into pivot assembly (3).
NOTE
When installing a new pivot assembly (3),
burnishing the 2-100" I.D. bore may be
required to facilitate assembly of landing
gear strut (5).
Bore Diameter
Piston Rod Diameter .....
Piston Stroke
...........
in
.998 in.
2.970 in
a. Remove screw (23). Remove end gland (22) by
unscrewing end gland from cylinder body (15).
b. Remove end cap (6). Remove AN3164R nuts
(9) if installed and remove cap (5) by pulling from
cylinder body (15). Using a small rod, push piston
(18) from cylinder body (15).
c. Remove cap (5) from shaft (14) by removing
retainer (2) and washer (3).
d. Remove shaft (14), sector (12) and washer (11)
from cylinder body (15).
e. Remove setscrew (13) from sector (12).
Remove section from shaft (14).
NOTE
Unless defective, do not remove name
plate, bearing (7) and (10) or roller (8).
c. Align hole in plug (20) with holes in pivot assembly
(3) using special tool No. SE934.
NOTE
Special tool No. SE934 is available from
Cessna Parts Distribution (CPD 2) through
Cessna Service Stations. This tool is designed
to install strut attaching bolt without
damaging the O-rings in the plug.
d. Install the strut attaching bolt (31) by pushing the
SE934 tool through the aligned holes of the pivot
I assembly (3), strut (5), and plug (20), with the threaded
nut and washer
end
the
bolt
(31). Install
end of
of
the
bolt
(31).
Install and
and tighten
tighten nut and washer
on the
the bolt (31).
system in accordance with
e. Fill and bleed brake system in accordance with
paragraph 5-77 in this manual.
5-9. MAIN LANDING GEAR ACTUATOR.
f. Remove O-ring (17) and back-up ring (16) from
cylinder body (15). Discard O-ring (17).
g. Remove O-ring (20) and back-ring (21) from
end gland (22). Discard O-ring (20).
h. Remove and discard O-ring (19) from piston
(18).
5-12.
a.
INSPECTION OF PARTS.
Thoroughly clean all parts in cleaning solvent
(Federal Specification PS-661. or equivalent.)
b
Inspect all threaded surfaces for cleanliness
cracks(5),
washers (3) and (11), sector
c. Inspect cap (5), washers (3) and (11), sector
(12), shaft (14), piston (18), roller (8). if removed.
and cylinder body (15) for cracks, chips, scratches.
scoring, wear or surface irregularities which may
affect their function or the overall operation of the
actuator.
d. Inspect bearings (7) and (10). if removed, for
Revision 3
5-13
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Lubricate sector, piston, rack gears and
all bearings with MIL-G-21164 lubricant
during assembly of the main gear actuator.
\
NOTE
22
21
Install new packings, lubricated with a film
of Petrolatum W-P-236, hydraulic fluid MILH-5606, or Dow-Corning DC-7.
1
19
^>^
18
1. Bolt
20.
21.
22.
23.
Figure 5-2.
O-Ring
Back-Up Ring
End Gland
Screw
Main Landing Gear Actuator Assembly
freedom of motion, scores, scratches or Brinnel
marks.
5-13. PARTS REPAIR/REPLACEMENT. Repair
of small parts of the main landing gear actuator is
impractical Replace all defective parts. Minor
scratches or score marks may be removed by
polishing with abrasive crocus cloth (Federal
Specification P-C-458), providing their removal
does not affect operation of the unit. During
assembly, install all new packings.
5-14. MAIN GEAR ACTUATOR REASSEMBLY.
(Refer to figure 5-2.)
NOTE
Use MIL-G-2116C lubricant on roller (8),
bearings (7) and (10), if removed, and
sector (12) when installing in cylinder
body (15).
a. It bearings (7) and roller (8) were removed,
press one bearing (7) into cylinder body (15) until it
is flush. Install roller (8) and press second bearing
(7) in place to hold roller. Use care to prevent
damage to bearings and roller.
b. If bearing (10) was removed, press bearing
into cap (5) until flush.
c. Assemble sector (12) on shaft (14). aligning
index marks on shaft and sector. Install setscrew
(13), making sure that setscrew enters shaft.
d. Position washer (11) and cap (5) on shaft (14).
Install washer (3) and retainer (2) on shaft.
e. If actuator is to be installed in aircraft. install
cap and shaft assembly on cylinder body with bolts
(1) and washers (4). If actuator is not to be
installed in aircraft, install cap and shaft assembly
on cylinder body with bolts (1). washers (4) and
AN316-4R nuts (9).
f. Install back-up ring (16) and O-ring (17) in
cylinder body bore. Install new O-ring (19) on
piston (18).
NOTE
Install new packings, lubricated with a film
of Petrolatum W-P-236, hydraulic fluid MILH-5606, or Dow-Corning DC-7.
g. Rotate shaft (14) so that teeth on sector (12) are
toward cylinder body.
5-15
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
c. Connect hydraulic lines to actuator.
d. Install access plates on bulkhead forward of
actuator.
e. Connect brake line at wheel cylinder. Fill and
bleed brake system in accordance with instructions
in applicable paragraph in this Section.
f. Rig landing gear in accordance with
procedures outlined in applicable paragraph in this
Section.
g. Remove aircraft from jacks and install access
covers, carpeting and seats removed for access.
5-16. MAIN LANDING GEAR STRUT-TOACTUATOR LINKAGE. (Refer to figure 5-1.)
5-7. DESCRIPTION. Each main landing gear
actuator attaches directly to a pivot assembly,
which in turn is attached to. and rotates its own
main landing gear strut.
5-18. PIVOT ASSEMBLY REMOVAL. (Refer to
figure 5-1.)
a. Remove main landing gear strut as outlined in
paragraph 5-7.
b. Loosen nut (12) and telescope pivot shaft (13)
inboard to free pivot assembly (3) from bearing (6)
in inboard support (2).
c. Remove pivot assembly (3), bearing (8)
and bearing race (7).
5-19. PIVOT ASSEMBLY INSTALLATION.
(Refer to figure 5-1.)
a. Install bearing (8) and race (7) on shaft of
pivot assembly (3); install tab washer (11) and nut
(12) on pivot shaft (13).
b. Position shaft of pivot assembly (3) into
bearing (6) in inboard support (2). Lubricate
bearing (6) with MIL-G-21164 grease. Be sure
thrust bearing and race are correctly positioned.
c. Telescope_pivot shaft (13) and fit shaft (13) into
- bushing (16) in outboard support (4).
d. Tighten nut (12) firmly and safety in place,
bending corresponding tangs of washer (11). Pivot
assembly shall rotate freely.
5-20. MAIN GEAR UPLOCK MECHANISM.
(Refer to figure 5-4.)
5-21. DESCRIPTION. The uplock actuator
cylinder and latches for the main landing gear are
located on the aft side of canted bulkhead station
106.00 (refer to Section 1 of this manual.) The
latches are controlled by a single actuator, located
on the aircraft centerline, by means of bellcrank
and linkage assemblies.
WARNING
Before working in landing gear wheel
wells, PULL-OFF hydraulic pump circult breakers. Thru Serial 21062273,
the pump circuit breaker is located in
the circuit breaker panel, located immediately forward of the pilot's control
wheel. Beginning with Serial 21062274,
the pump circuit breaker is located in
the circuit breaker panel, located immediately forward of the left forward
doorpost. The hydro-electric power
pack system is designed to pressurize
the landing gear DOOR CLOSE system
to 1500 psi at any time the master switch
is turned on. Injury might occur to
someone working in wheel well area if
master switch is turned on for any
reason.
a. Turn master switch OFF and. using hand
pump, open landing gear doors.
b. Components of the main landing gear uplock
system are readily accessible on the aft side of
canted bulkhead station 106.00 (refer to Section 1 of
this manual.)
c. Components may be removed or installed
using figure 5-4 as a guide.
d. Upon installation, rig uplocks in accordance
with applicable paragraph in this Section.
5-22A.
UPLOCK ACTUATOR.
5-23. UPLOCK ACTUATOR DISASSEMBLY.
(Refer to figure 5-5.)
NOTE
Leading particulars of the actuators
Cylinder Bore Diameter . .
0.749 +.002.-.000 in.
Piston Diameter .
. .
. . 0.747+.000.-.001 in.
Stroke (to unseat valve) .
.
0.719 ± .031 in.
a. Remove fitting (5), spring (7) and balls (8) and
(9).
b. Cut safety wire and unscrew end plug (19)
from barrel and valve body (12).
c. If end fitting (1) is installed, loosen nut (2) and
remove end fitting from barrel and valve body.
d. Remove springs (18) and (17) and push piston
and rod (13) from barrel and valve body.
5-22. REMOVAL AND INSTALLATION OF MAIN
GEAR UPLOCK MECHANISM. (Refer to figure 54.)
5-17
MODEL 210 & T210 SERIES SERVICE MANUAL
CANTED BULKHEAD STA. 106.00
(REFER TO FIGURE 1-1 FOR
MODEL 210 &T210 SERIES SERVICE MANUAL
10
10
2
Figure 5-6A.
Nut
End Fitting
Body
Spring
Fitting
Spring
Ball
8. Ball
9. Piston/Rod
10. O-Ring
11. Back-up Ring
12. Jamnut
13. Rod End
Main Landing Gear Downlock Actuator.
rig the main landing gear in accordance with
procedures outlined in the applicable paragraph in
this Section.
5-28A.
1.
2.
3.
4.
5.
6.
7.
DOWNLOCK ACTUATOR.
5-29.
DISASSEMBLY. (Refer to figure 5-6A.)
a. Loosennut (1) and unscrew end fitting (2) from body
(3). Spring (4) can also be removed.
b. Remove fitting (5), spring (6), ball (7), and ball (8)
from body (3).
c. Remove piston/rod (9) from body (3).
d. Remove and discard all packings and back-up rings
from end fitting (2), body (3), and piston/rod (9).
5-29A. INSPECTION AND REPAIR.
a. Inspect all threaded surfaces for cleanliness and for
freedom from cracks and excessive wear.
b. Inspect spring (6) for evidence of breaks and
"ia.
*~ ~~~~~~~~distortion,~
c. Inspect piston spring (4) for evidence of breaks and
distortion.
d. Inspect end fitting, piston/rod, barrel, valve
body, balls and ball seats for cracks, scratches,
scoring, wear or surface irregularities which might
affect their function or the overall function of the unit.
e. Repair of most parts of the unlock actuator is
impractical. Replace defective parts. Minor
scratches and scores may be removed by polishing
with fine abrasive crocus cloth (Federal Secification
PC-458), providing their removal does not affecc
operation of the unit.
5-29B.
REASSEMBLY.
NOTE
Install new O-rings and back-up rrins lubricated with a film of Petrolatum W-P-23 6
hydraulic fluid MIL-H-5606, or Dow-Corning DC-7.
Assemble by reversing procedures outlin.ed in
paragraph 5-29.
Revision 3
5-20A/5-20B Blank
MODEL 210 & T210 SERIES SERVICE MANUAL
5-30.
MAIN LANDING GEAR DOOR SYSTEM.
c. Repair of most parts of the landing gear door
actuator assembly is impractical. Replace
defective parts with new parts.
d. Minor scratches may be removed by polishing
with fine abrasive crocus cloth (Federal
Specification PC-458), providing their removal does
not affect the operation of the unit.
5-31. DESCRIPTION. Main gear doors open for
main gear retraction or extension and return to
closed positions at the close of either cycle. The
strut doors are opened and closed by a doubleacting hydraulic actuator. The wheel doors are
actuated by a double-actuating hydraulic actuator
for each door. The actuators are held closed by the
door close system accumulator.
5-32.
5-33C.
REMOVAL AND INSTALLATION OF MAIN
Petrolatum
with Petrolatum
VV-P-236. hydraulic
fluid MIL-H-5606. or Dow Cornin DC-7
a. Install new O-ring and back-up ring in gland
and install gland on piston rod. Use care to
prevent damage to O-rings and back-up rings.
b. Install new 0-rings and back-up rings on
piston and on gland.
c. Install piston rod and gland into cylinder and
install retaining ring. Use care to prevent damage
to O-rings and back-up rings.
WARNING
a. Peel back carpet as required and remove access
cover in center of floorboard just forward of rear
seat.
b. Open doors using hand pump then disconnect
hydraulic lines at actuator. Cap or plug lines and
fittings.
c. Remove bolts at each end of actuator attaching
rod end to bellcrank and actuator body to mounting
bracket. Remove actuator from aircraft.
d. Reverse procedure to install actuator.
5-33A. DISASSEMBLY. (Refer to figure 5-8.)
b. Remove retaining ring (1) from end of cylinder
(6).
c. Pull piston rod (5), end gland ( 4 ) from
cylinder (6). A sharp blast of air applied to the
hydraulic port at bearing end of cylinder may be
used to remove piston rod.
d. Remove end gland (4 ) from piston rod (5).
rings from gland and piston rod.
5-33B. INSPECTION. (Refer to figure 5-8.)
a. Inspect all threaded surfaces for cleanliness
and for freedom of cracks and excessive wear or
damage.
b. Inspect end gland ( 4 ), piston rod (5) and
cylinder (6) for cracks, chips, scratches, scoring.
wear or surface irregularities which might affect
their function or the overall function of the door
actuator.
hydraulic
during assembly.
5-33.
MAIN GEAR STRUT DOOR ACTUATOR
REMOVAL AND INSTALLATION.
circuit breaker before disconnecting any hydraulic lines in the landing gear system.
(Refer to figure 5-8.)
NOTE
Lubricate all O-rings and back-up rings
5-32. REMOVAL AND INSTALLATION OF MAIN
GEAR STRUT AND WHEEL DOORS. (Refer to figure
5-7.)
5*)
a. Open landing gear doors.
b. Disconnect door from actuator linkage by
removing pin or bolt.
c. Remove door hinge pins or bolts.
d. Install door by reversing the preceding steps.
e. Rig doors in accordance with applicable
paragraph.
Turn master switch "off" and pull pump motor
REASSEMBLY.
the
5-34. MAIN WHEEL DOOR ACTUATOR REMOVAL.
a. Open landing gear doors.
b. Disconnect and cap or plug hydraulic hoses at
actuator
Disconnect
actuator rod by removing
attaching
attaching nut
nut and
and bolt
bolt at
at door.
door.
d. Remove nut and bolt attaching actuator to
fuselage bracket and remove actuator.
5-35. MAIN WHEEL DOOR ACTUATOR DISASSEMBLY. (Thru Serial 21062273, refer to figure 5-8A.)
a. Loosen check nut (2) and remove rod end (1).
b. Remove retaining ring (3) from end of cylinder
(10).
o. Pull piston rod (8). gland (6) or (7) from
cylinder (10). A sharp blast of air applied to the
hydraulic port at bearing end of cylinder may be
used to remove piston rod.
d. Remove gland (6) or (7) from piston rod (8).
e. Remove and discard back-up rings and Orings from gland and piston rod.
f. Do not remove bearing (9) unless it is
defective.
5-35A INSPECTION. (ThruSerial21062273.)
a. Inspect all threaded surfaces for cleanliness
and for freedom of cracks and excessive wear or
b. Inspect gland (6) or (7). piston rod (8) and
cylinder (10) for cracks, chips, scratches. scoring,
wear or surface irregularities which might affect
their function or the overall function of the door
actuator.
5-21
MODEL 210 & T210 SERIES SERVICE MANUAL
c. Repair of most parts of the landing gear door
actuator assembly is impractical. Replace
defective parts with new parts.
d. Minor scratches may be removed by polishing
with fine abrasive crocus cloth (Federal
Specification PC-458), providing their removal does
not affect the operation of the unit.
5-35B. REASSEMBLY.
to figure 5-8A.)
(Thru Serial 21062273.) (Refer
NOTE
Lubricate all O-rings and back-up rings with
a film of Petrolatum VV-P-236, Hydraulic
fluid MIL-H-5606, or Dow-Corning DC-7.
a. Install new O-ring and back-up ring in gland
and install gland on piston rod. Use care to
prevent damage to O-rings and back-up rings.
b. Install new O-rings and back-up rings on
piston and on gland.
c. Install piston rod and gland into cylinder and
install retaining ring. Use care to prevent damage
to O-rings and back-up rings.
d. Install lock nut and rod end.
NOTE
If bearing (9) was removed, install and
stake six places, three on each side.
5-36. MAIN WHEEL DOOR ACTUATOR DISASSEMBLY. (Beginning with Serial 21062274, refer to
figure 5-8A.)
a. Loosen check nut (2) and remove rod end (1).
b. Remove safety wire from end fitting (11) unscrew end fitting from actuator cylinder (10).
c. Pull piston rod (8) from cylinder.
d. Remove and discard back-up rings and O-rings
from end fitting and piston rod.
e. Do not remove bearing (9) unless it is defective.
5-36A. INSPECTION. (Beginning with Serial
21062274. )
a. Inspect all threaded surfaces for cleanliness
and for freedom of cracks and excessive wear or
dammage.
b. Inspect end fitting, piston rod and cylinder for
cracks, chips, scratches, scoring, wear or surface
irregularities which might affect their function or
the overall function of the door actuator.
c. Repair of most parts of the gear door actuator
is impractical. Replace defective parts with new
parts.
d. Minor scratches may be removed by polishing
with fine abrasive crocus cloth (Federal Specification PC-458. providing their removal does not
affect the operation of the unit.
5-36B. REASSEMBLY. (Beginning with Serial
21062274.)
a. Install new O-ring and back-up ring inside end
fitting. Install new O-ring on outside of end fitting.
b. Install new O-rings and back-up rings on
piston.
c. Install piston in cylinder using care to avoid
dammaging O-rings and back-up rings.
d. Install end fitting on piston rod and screw into
cylinder. Use care to prevent dammage to O-ring
and back-up ring inside end fitting.
e. Tighten end fitting and install new safety wire.
NOTE
If bearing (9) was removed, install and
stake six places. three on each side.
5-37.
MAIN WHEEL AND TIRE ASSEMBLY.
5-38. DESCRIPTION. The aircraft may be
equipped with either Cleveland or McCauley wheel
and tire assemblies. Separate disassembly,
inspection and reassembly instructions are
provided for each type.
CAUTION
Use of recapped tires or new tires not listed
on the aircraft equipment list are not recommended due to possible interference between
the tire and structure when landing gear is
in the retracted position.
REMOVAL OF MAIN WHEEL AND TIRE
5-39.
ASSEMBLY. (Refer to figure 5-1.)
It is not necessary to remove the main wheel
to reline brakes or remove brake parts, other
than the brake disc or torque plate.
a. Using thejack point under step on main gear strut,
jack up wheel being removed in accordance with
procedures outlined in Section 2.
b. Remove hub caps (25).
c. Remove cotter pin (32) and nut (24).
d. Remove bolts and washers attaching back plate, and
remove back plate (Index 22, figure 5-9, Sheet 1).
e. Pull wheel and tire assembly (23) from axle (21).
Revision 3
5-22A
MODEL 210 & T210 SERIES SERVICE MANUAL
2
5-22B
MODEL 210 & T210 SERIES SERVICE MANUAL
coating has been removed, the area should be cleaned
| thoroughly, primed with nonzinc chromate, and
repainted with aluminum lacquer.
c. Brake disc should be replaced if excessively
scored or warped. Small nicks and scratches should
be sanded smooth. See paragraph 5-72.
d. Bearing cups and cones should be inspected
carefully for damage and discoloration. After
cleaning. repack cones with clean aircraft wheel
bearing grease (Section 2) before installation in the
wheel.
5-42. REASSEMBLY OF CLEVELAND MAIN
WHEEL AND TIRE ASSEMBLY. (Refer to figure
5-9.)
a. Insert thru-bolts through brake disc and
position in the inner wheel half. using the bolts to
guide the disc. Assure the disc is bottomed in
wheel half.
b. Position tire and tube with the inflation valve
through hole in outboard wheel half. Place inner
wheel half in position. Apply a light force to bring
wheel halves together. Maintaining the light force,
assemble a washer and nut on one thru-bolt and
tighten snugly. Assemble remaining washers and
nuts on thru-bolts and torque to 150 lb-in.
CAUTION
Uneven or improper torque of thru-bolt
nuts may cause failure of bolts, with
resultant wheel failure.
c. Clean and repack bearing cones with clean
aircraft wheel bearing grease (refer to Section 2 of
this manual).
d. Assemble bearing cones. grease seal felts and
rings into wheel halves.
e. Inflate tire to seat tire beads, then adjust to correct
| pressure specified in figure 1-1.
DISASSEMBLY OF MCCAULEY TWO-PIECE
5-43.
EASSEMBLY. (Referto
*MAIN WHEELAND TIRE
figure5-9,Sheet2.)
a. Deflate tire and break tire beads loose.
CAUTION
CAUTION
Avoid damaging wheel flange when breaking
tire beads loose. A scratch, gouge, or nick
may cause wheel failure.
b. Remove thru-bolts (24) and separate wheel halves (6)
and (10), removing tire (8), tube (9), and brake disc (13).
c. Remove grease seal retainers (2) and (4), grease seal
felts (3), and bearing cones (5) from wheel halves (6) and
(10).
5-26
Revision 3
NOTE
The bearing cups are a press fit in the
wheel halves and should not be removed
unless replacement is necessary. To
remove the bearing cups. heat the wheel
half in boiling water for 15 minutes.
Using an arbor press. if available, press
out the bearing cup and press in the new
cup while the wheel is still hot.
5-44. INSPECTION AND REPAIR OF McCAULEY
TWO-PIECE MAIN WHEEL AND TIRE ASSEMBLY
( Refer to figure 5-9. )
a. Clean all metal parts and the grease seal felts
in solvent and dry thoroughly.
b. Inspect wheel halves for cracks. Cracked
wheel halves should be replaced. Sand out nicks.
gouges and corroded areas. When the protective
coating has been removed, the area should be cleaned
thoroughly, primed with nonzinc chromate, and
repainted with aluminum lacquer.
c. Brake disc should be replaced if excessively
scored or warped. Small nicks and scratches should
be sanded smooth. See paragraph 5-72.
d. Bearing cups and cones should be inspected
carefully for damage and discoloration. After
cleaning, repack cones with clean aircraft wheel
bearing grease (Section 2) before installation in the
wheel.
5-44A. REASSEMBLY OF McCAULEY TWO-PIECE
MAIN WHEEL AND TIRE ASSEMBLY. (Refer to
figure 5-9.)
a. Insert thru-bolts through brake disc and
position in the inner wheel half. using the bolts to
guide the disc. Assure the disc is bottomed in
wheel half.
b. Position tire and tube with the inflation valve
through hole in outboard wheel half. Place inner
wheel half in position. Apply a light force to bring
wheel halves together. Maintaining the light force.
assemble a washer and nut on one thru-bolt and
tighten snugly. Assemble remaining washers and
nuts on thru-bolts-and
torque
to 150
150 lb-in.
lb-in
and
torque to
thru bolts
CAUTION
Uneven or improper torque of thru-bolt
nuts may cause failure of bolts. with
resultant wheel failure.
c. Clean and repack bearing cones with clean
aircraft wheel bearing grease (refer to Section 2 of
this manual).
d. Assemble bearing cones (5), grease seal felts (3), and
grease seal retainers (2) and (4) into wheel halves (6) and
(10).
e. Inflate tire to seat tire beads, then adjust to correct
pressure.
|
MODEL 210 & T210 SERIES SERVICE MANUAL
Injury can result from attempting to
remove wheel flanges with the tire and
tube inflated. Avoid damaging wheel
flanges when breaking tire beads loose.
c. Sand out smooth nicks, gouges and corroded
areas. When the protective coating has been
removed, the area should be cleaned thoroughly.
primed with zinc chromate and painted with
aluminum lacquer.
d. Brake disc should be replaced if excessively
scored or warped. Small nicks and scratches should
be sanded smooth. See paragraph 5-72.
e.
Carefully inspect bearing cones and cups for
damage and discoloration. After cleaning, pack
bearing grease (refer to Section 2 of this manual)
before installing in the wheel hub.
A scratch, gouge or nick in wheel flanges
5-45B. REASSEMBLY OF McCAULEY THREE PIECE
Remove valve core and deflate tire and
MAIN WHEEL AND TIRE ASSEMBLY. (Refer to figure
5-9. )
5-45. DISASSEMBLY OF McCAULEY THREE PIECE
MAIN WHEEL AND TIRE ASSEMBLY. (Refer to
figure 5-9. )
a. Remove screws attaching hub cap; remove hub
cap.
WARNING
b.
~
c. Remove cap screws
d. Remove brake disc.
e. Separate wheel flanges from wheel hub.
cones. on each side of wheel hub.
bearing
Retain spacers
d. Remove wheel hub from tire.
rings
retainer
and remove
g. Remove
grease seal retainers, grease seal felts and
NOTE
The bearing cups (races) are a press fit in
the wheel hub and should not be removed
unless a new part is to be installed. To
remove the bearing cup, heat wheel hub
in boiling water for 30 minutes, or in an
oven not to exceed 121°C (250°F). Using
an arbor press, if available, press out the
bearing cup and press in the new bearing
cup while the wheel is still hot.
5-45A. INSPECTION AND REPAIR OF McCAULEY
THREE PIECE MAIN WHEEL AND TIRE ASSEMBLY,
(Refer to figure 5-9. )
a. Clean all metal parts, grease seal felts and
solvent and dry
cleaning
in
phenolic spacers
thoroughly.
b. Inspect wheel flanges and wheel hub for
cracks. Cracked wheel flanges or hub shall be
discarded and new parts installed.
SHOP NOTES:
5-26B
b. Place spacer and wheel flange on inboard side
of wheel hub (opposite of tube inflation stem), then
place washer under head of each capscrew and
start capscrew into hub threads.
c. Place spacer and wheel flange on other side
and align valve stem in cutout in wheel Range..
d. Place washer under head of each capscrew and
start capscrews into hub threads.
CAUTION
Be sure that spacers and wheel flanges
are seated on flanges of wheel hub.
Uneven or improper torque of capscrews
can cause failure of the capscrews or hub
threads with resultant wheel failure.
e. Tighten capscrews evenly and torque to 190200 lb in.
f. Clean and pack bearing cones with clean
aircraft wheel bearing grease. Refer to Section 2 of
this manual for grease type.
g. Assemble bearing cones, grease seal felts and
retainer into wheel hub.
h. Inflate tire to seat tire beads, then adjust to
correct pressure specified in figure 1-1.
5-46 INSTALLATION OF MAIN WHEEL AND
TIRE ASSEMBLY.
MODEL 210 & T210 SERIES SERVICE MANUAL
a.
Place wheel on axle.
b. Install axle nut and tighten until a slight
bearing drag is obvious when the wheel is rotated.
Back off nut to nearest castellation and install
cotter pin.
c. Place brake back plate in position and secure
with bolts and washers. Safety wire the bolts.
d. Install hub caps.
5-47. MAIN WHEEL DOOR CLOSE SYSTEM
ACCUMULATOR. (Refer to figure 5-10.)
5-48. DESCRIPTION. The accumulator serves
two purposes. This unit maintains pressure in the
door-close system, keeping the main wheel doors
up and closed. The accumulator also dampens
pressure surge and serves as a reservoir to offset
|normal leak-down in the system.
WARNING
WARNING
BEFORE WORKING IN WHEEL WELL
AREA, PULL HYDRAULIC PUMP
CIRCUIT BREAKER OFF.
5-49. REMOVAL OF ACCUMULATOR.
figure 5-10.)
(Refer to
WARNING
Filler and safety valve (8) does not contain a
core. To release accumulator pressure, loosen
nut on end of valve. If the valve installed
contains a core, the valve should be replaced
with a valve which does not contain a core.
Injury can occur if pressure is not released
properly.
a. Open main gear doors. This will drop
hydraulic pressure to zero.
b. Relieve accumulator pressure by turning nut
on end of valve approximately 1/4 turn.
c. Disconnect and plug or cap hydraulic line at
accumulator.
d. Remove bolt, washer, spacer and nut at
outboard end and remove clamp, screw and nut at
inboard end; remove accumulator.
DISASSEMBLY AND REASSEMBLY OF
5-50.
ACCUMULATOR. (Refer to figure 5-10.)
a. Remove retainer (18) only after ensuring that
pressure has been relieved. Remove gland (19), piston
| (20), and filler and safety valve (8) if required.
b. Remove and discard packings (22) and back-up
rings (23).
c. Reverse the preceding steps, using new
packings and back-up rings, for reassembly of the
accumulator.
~~.050
~~~~~NOTE
Install new packings and back-up rings lubricated with a film of Petrolatum VV-P236, hydraulic fluid MIL-H-5606, or DowCorning DC-7.
5-51. INSTALLATION OF ACCUMULATOR.
(Refer to figure 5-10.)
WARNING
BEFORE WORKING IN WHEEL WELL
AREA. PULL HYDRAULIC PUMP
CIRCUIT BREAKER OFF.
a. Install bolt, washer, spacer and nut at
outboard end and clamp screw and nut at inboard
end.
b. Connect hydraulic line at accumulator.
c. Pressurize accumulator with nitrogen or dry
air to 500 + 50 psig. Hydraulic pressure should be
zero.
NOTE
Adapter hose and fitting kit (nitrogen bottle
to accumulator) number ZN216, available
from Cessna Parts Distribution tCPD 2)
through Cessna Service Stations, can be used
to charge the accumulator.
5-52. MAIN WHEEL AND AXLE REMOVAL.
(Refer to figure 5-1.)
a. Remove hub caps.
b. Remove wheel from axle in accordance with
procedures outlined in paragraph 5-39.
c. Disconnect, drain and plug hydraulic brake
line at the brake cylinder.
d. Remove bolts, washers, nuts and stud secruing
axle and brake components to fitting at lower end
of strut.
NOTE
When removing axle from strut fitting,
note number and position of wheel
alignment shim. Mark these shims or
tape together carefully so they can be
reinstalled in exactly the same position to
ensure that wheel alignment is not
disturbed. Also. note position of stud
attaching axle to fitting so that the stud
may be installed in the same position.
Stud is the uplock for the main gear.
5-53. MAIN WHEEL AND AXLE
INSTALLATION. (Refer to figure 5-1.)
a. Secure axle and brake components to strut
fitting, making sure that wheel alignment shims
and stud are reinstalled in their original position.
NOTE
Shim: P/N 1241061-3, available from Cessna
Parts Distribution (CPD 2) through Cessna
Service Stations, can be installed between
axle and fitting, if necessary, to maintain
inch minimum clearance between axl
fitting and brake disc.
b. Install wheel assembly on axle in accordance
with paragraph 5-46.
c. Connect hydraulic brake line to brake
cylinder.
d. Fill and bleed affected brake system-.
e. Install hub caps.
f. Check wheel alignment.
Revision 3
5-27
MODEL 210 & T210 SERIES SERVICE MANUAL
5-54. MAIN WHEEL ALIGNMENT. Correct main
wheel alignment is obtained through the use of
tapered shims between the landing gear strut and
the flange of the axle. Refer to figure 5-11 for
procedures to use in checking alignment. Wheel
shims. and the correction imposed on the wheel by
the various shims, are listed in the illustration.
NOTE
Failure to obtain acceptable wheel
alignment through the use of the shims
indicates a deformed main gear strut or a
bent axle.
5-55. WHEEL BALANCING. Since uneven tire wear
is usually the cause of wheel unbalance, replacing the
tire probably will correct this condition. Tire and
5-58.
tube manufacturing tolerances permit a specified amount
of static unbalance. The lightweight point of the tire is
marked with a red dot on the tire sidewall, and the
heavyweight point of the tube is marked with a
contrasting color line (usually near the valve stem).
When installing a new tire, place these marks adjacent to
each other. If a wheel becomes unbalanced during
service, it may be statically balanced. Wheel balancing
equipment is available from Cessna Parts Distribution
(CPD 2) through Cessna Service Stations.
5-56.
BRAKE SYSTEM.
5-57. DESCRIPTION. The hydraulic brake system
consists of two master cylinders, brake lines, connecting each master cylinder to its corresponding
wheel brake cylinder, and the single, disc-type brake
assembly, located at each main landing gear wheel.
TROUBLE SHOOTING.
TROUBLE
DRAGGING BRAKES.
BRAKES FAIL TO OPERATE.
PROBABLE CAUSE
REMEDY
Brake pedal binding.
Check and adjust properly.
Parking brake linkage holding
brake pedal down.
Check and adjust properly.
Worn or broken piston return
spring. (In master cylinder.)
Repair or replace master
cylinder.
Insufficient clearance at LockO-Seal in master cylinder.
Adjust as shown in figure 5-12.
Restriction in hydraulic lines
or restriction in compensating
oort in master brake cylinder.
Drain brake lines and clear the
inside of the brake line with filtered compressed air. Fill and
bleed brakes. If cleaning the
lines fail to give satisfactorY
results, the master cylinder may
be faulty and should be repaired.
Worn, scored, or warped brake
discs.
Replace brake disc and linings.
Damage or accumulated dirt
restricting free movement
of wheel brake parts.
Clean and repair or replace parts
as necessary.
Leak in system.
Check entire system for leaks
If brake master cylinders or
wheel assemblies are leaking, they
should be repaired or replaced.
Air in system.
Bleed system.
Lack of fluid in master cylinders.
Fill and bleed systems.
Revision 3
5-29
MODEL 210 & T210 SERIES SERVICE MANUAL
brake cylinders.
b. Remove front seats and rudder bar shield
for access to brake master cylinders.
c. Disconnect parking brake linkage and
disconnect brake master cylinders from
rudder pedals.
d. Disconnect hydraulic hose from brake
master cylinders and remove cylinders.
e. Plug or cap hydraulic fittings, hose and
lines to prevent entry of foreign material.
hole is open.
j.
Install setscrew (5).
5-62.
seats.
BRAKE MASTER CYLINDER
DISASSEMBLY. (Refer to figure 5-12.)
a. Unscrew clevis (1) and jamb nut (2).
b. Remove screw (18).
c. Remove filler plug (17) and setscrew (5).
d. Unscrew cover (4) and remove up over
piston rod (3).
e. Remove piston rod (3) and compensating
sleeve (16).
f. Slide sleeve (16) up over rod (3).
g. Unscrew nut (12) from threads of piston rod
(3).
h. Remove piston spring (13) and O-ring (9) from
piston (14).
5-63. BRAKE MASTER CYLINDER
INSPECTION AND REPAIR. (Refer to figure
5-12.) Repair is limited to installation of new
parts, cleaning and adjusting. (Refer to
reassembly paragraph for adjustment.) Use
clean hydraulic fluid (MIL-H-5606) as a
lubricant during reassembly of the
cylinders. Inspect Lock-O-Seal (Parker Seal
Co. P/N 800-001-6) and replace if damaged.
Replace all O-rings. Filler plug must be
vented so pressure cannot build up in the
reservoir during brake operation. Remove
plug and drill 1/16-inch hole, 30 ° from
vertical, if plug is not vented.
5-64. BRAKE MASTER CYLINDER
REASSEMBLY. (Refer to figure 5-12.)
a. Install Lock-O-Seal (15) at bottom of
piston rod (3).
b. Install O-ring (9) in groove in piston
(14); insert piston spring (13) into piston, and
slide assembly up on bottom threaded portion
of piston rod (3).
c. Run nut (12) up threads to spring (13):
Tighten nut enough to obtain 0.040 ± 0.005-inch
clearance between top of piston and bottom of
Lock-O-Seal, as shown in the figure.
d. Install piston return spring (11) into
cylinder (10) portion of body (7).
e. Install piston rod (3) through spring (11).
f. Slide compensating sleeve (16) over rod
(3).
g. Install cover (4) and screw (18).
h. Install
Install jamb
jamb nut
nut.
and clevis (1)
i. Install filler plug (17), making sure vent
5-65. BRAKE MASTER CYLINDER
INSTALLATION.
a. Connect hydraulic hoses to brake master
cylinders and install cylinders
b. Connect brake master cylinders to
rudder pedals and connect parking brake
linkage.
c. Install rudder bar shield and install front
d. Install bleeder screw at wheel brake
assembly and fill and bleed brake system in
accordance with applicable paragraph in this
Section.
5-66.
HYDRAULIC BRAKE LINES.
5-67. DESCRIPTION. The brake lines are of
rigid tubing, except for flexible hose used at
the brake master cylinders. A separate line
is used to connect each brake master cylinder
to its corresponding wheel brake cylinder.
WARNING
After connecting brake hose, ensure that
hose does not contact or rub against
brake disc, causing brake hose failure.
5-68. WHEEL BRAKE ASSEMBLIES.
(Refer to figure 5-9.)
5-69. DESCRIPTION. The wheel brake
assemblies employ a floating brake assembly
and a disc which is attached to the main
wheel.
5-70. WHEEL BRAKE REMOVAL. (Refer to
figure 5-9.) Wheel brake assemblies can be
removed by disconnecting the brake line
(drain fluid when disconnecting line) and
removing the brake back plate. The brake
disc is removed after the wheel is removed
and disassembled. -To remove the torque
plate, remove wheel and axle.
5-71.
Refer
brake
guide
WHEEL BRAKE DISASSEMBLY.
to figure 5-9 for a breakdown of wheel
parts. This figure may be used as a
for disassembling the wheel brakes.
5-72. WHEEL BRAKE INSPECTION AND
REPAIR.
a. Clean all parts except brake linings and
O-rings in dry cleaning solvent and dry
thoroughly.
b.ll
new O-rings. If O-ring reuse
is necessary, wipe with a clean cloth
saturated in hydraulic fluid and inspect for
damage.
5-32
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Thorough cleaning is important. Dirt
and chips are the greatest single cause of
malfunctions in the hydraulic brake
system.
c.
Check brake lining for deterioration and
paragraph for maximum wear limit.)
d. Inspect brake cylinder bore for scoring. A
scored cylinder will leak or cause rapid O-ring
wear. Install a new-brake cylinder if the bore is
scored,
e. If the anchor bolts of the brake assembly are
nicked or gouged, they shall be sanded smooth to
prevent binding with the pressure plate or torque
plate. When new anchor bolts are to be installed.
press out old bolts and install new bolts with a soft
mallet.
f. Inspect wheel brake disc for minimum thickness. If
disc is below minimum thickness, install a new part.
Minimum thicknesses are as follows:
Cleveland disc no. 164-15A: .340-inch
McCauley discs No. C30398 and C30615-3: .325-inch.
composition. Brake pads must be properly conditioned
(glazed) before use in order to provide optimum service
life. This is accomplished by a brake burn-in. Burn-in
also wears off brake high spots prior to operational use.
If brake use is required before burn-in, use brakes
intermittently at LOW taxi speeds.
5-74C. BRAKE BURN-IN.
CAUTION
Brake burn-in must be performed by a
qualified person familiar with acceleration
and stop distances of the airplane.
5-73. WHEEL BRAKE REASSEMBLY. (Refer to
figure 5-9.)
a. Non-asbestos Organic Composition Burn-in.
1. Taxi the airplane for 1500 feet, with engine at
1700 RPM, applying brake pedal force as need to
maintain 5 to 10 M.P.H. (5 to 9 Knots).
2. Allow brakes to cool for 10 to 15 minutes.
3. Apply brakes and check to see if a high
throttle static engine run-up can be held with normal
pedal force. If so, conditioning burn-in is complete.
4. If static run-up cannot be held, repeat Steps 1.
thru 3. as needed.
b. Metallic Composition Burn-in.
1. Taxi the airplane at 34 to 40 M.P.H. (30 to 35
Knots) and perform full stop braking application.
NOTE
CAUTION
Lubricate parts with a clean hydraulic
fluid during brake reassembly.
a. Refer to figure 5-9 as a guide while
reassembling wheel brakes.
5-74. WHEEL BRAKE INSTALLATION.
a. Place brake assembly in position with
pressure plate in place.application.
NOTE
If torque plate was removed. install as the
axle is installed, or install on axle. If the
brake disc was removed, install as wheel
is assembled.
Brake conditioning using successive stops at
higher speeds could cause brakes to overheat
resulting in warped discs and/or pressure
plates.
2. Without allowing brake discs to cool
substantially, repeat Step 1. for second full stop braking
3. Apply brakes and check to see if a high
throttle static engine run-up can be held with normal
pedal force. If so, conditioning burn-in is complete.
4. If static run-up cannot be held, repeat Steps 1.
thru 3. as needed.
NOTE
5-74A. BRAKE LININGS. (1977 THRU 1983
MODELS.) The pads are equipped with asbestos based
linings. When replacement is required, the new pads
must be properly conditioned (broken in) in order to
provide optimum service life. Conditioning will
generate sufficient heat to cure the resins in the
material, but will not cause the material to carburize
due to excessive heat. Condition the brakes by
performing a series of at least six light braking
applications from 25 to 40 MPH to a complete stop.
Allow the brake discs to partially cool after each stop.
Normal brake usage should generate enough
heat to maintain the glaze throughout the
life of the lining. Light brake usage can
cause the glaze to wear off, resulting in reduced brake performance. Visual inspection
of brake disc will indicate brake lining condition. A smooth, non-grooved surface indicates linings are properly glazed. Rough,
grooved linings must be reglazed. In such
cases, the lining may be conditioned again
following the instructions set forth above.
5-74B. NON-ASBESTOS ORGANIC OR METALLIC
BRAKE LININGS. Beginning with 1984 models, the
brake lining pads used in this assembly are either nonasbestos organic composition or iron based metallic
NOTE
5-34
Revision 3
Do not set parking brakes while brake discs
are hot.
MODEL 210 & T210 SERIES SERVICE MANUAL
5-75. CHECKING BRAKE LINING WEAR. New
brake lining should be installed when the existing
lining has worn to a thickness of 3/32-inch. A
3/32-inch strip of material held adjacent to each
lining can be used to determine amount of wear.
The shank end of a drill bit of the correct size can
also be used to determine wear of brake linings.
5-76. BRAKE LINING INSTALLATION. (Refer to
figure 5-9.)
a. Remove bolts securing back plate, and remove
back plate.
b. Pull brake cylinder out of torque plate and
slide pressure plate off anchor bolts.
c. Place back plate on a table with lining side
down flat. Center a 9/64-inch (or slightly smaller
punch in the rolled rivet, and hit the punch sharply
with a hammer. Punch out all rivets securing the
linings to the back plate in the same manner.
NOTE
A rivet setting kit, Part No. 199-1, is
available from Cessna Parts Distribution
(CPD 2) through Cessna Service Stations.
d. Clamp the flat side of the anvil in a vise.
e. Align new lining on back plate and place
brake rivet in hole with rivet head in the lining.
Place the head against the anvil.
f. Center rivet setting punch on lips of rivet.
While holding back plate down firmly against
lining, hit punch with a hmmer to set rivet. Repeat
blows on punch until lining is firmly against back
plate.
g. Realign the lining on the back plate and
install and set rivets in the 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 that anchor bolts slide
into the torque plate.
j. Install back plate with bolts and washers.
WARNING
After reinstallation of the brake assembly,
check brake line clearance to the disc in
the area above the axle.
5-77.
BRAKE SYSTEM BLEEDING.
NOTE
Bleeding with a clean hydraulic pressure
source connected to the wheel cylinder
bleeder is recommended.
a. Remove brake master cylinder filler plug and
screw flexible hose with appropriate fitting into the
filler hole at top of the brake master cylinder.
b. Immerse opposite end of flexible hose into a
container with enough hydraulic fluid to cover end
of the hose.
c. Connect a clean hydraulic pressure source,
such as a hydraulic hand pump or Hydro-Fill unit
to the bleeder valve in the wheel cylinder.
d. As fluid is pumped into the system. observe
the immersed end of the hose at the master cylinder
for evidence of air bubbles being forced from the
brake system. When bubbling has ceased, remove
bleeder source from wheel cylinder and tighten the
bleeder valve.
5-78. PARKING BRAKE SYSTEM. (Refer to
figure 5-13.)
5-79. DESCRIPTION. The parking brake system
consists of a handle and ratchet mechanism.
connected by a cable to linkage at the brake master
cylinders. Pulling out on the handle depresses
both brake master cylinder piston rods and the
handle ratchet locks the handle in this position
until the handle is turned and released.
5-80. REMOVAL AND INSTALLATION OF
COMPONENTS. Refer to figure 5-13 for relative location of system components. The
illustration may be used as a guide during removal and installation of components.
Revision 3
5-34A/(5-34B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
5-81. INSPECTION AND REPAIR OF SYSTEM
COMPONENTS. Inspect lines for leaks, cracks,
dents, chafing, improper radius, security,
corrosion, deterioration, obstructions and foreign
matter. Check brake master cylinders and repair
or replace as outlined in applicable paragraph in
this Section. Check parking brake handle and
ratchet for proper operation and release. Replace
worn or damaged parts.
5-82.
NOSE GEAR SYSTEM.
5-83. DESCRIPTION. The nose gear consists of a
pneudraulic shock strut assembly, mounted in a
trunnion assembly, a steering arm and bungee.
5-85.
shimmy dampener. uplock mechanism, nose wheel.
tire and tube, hub cap, bearings, seals and a doubleacting hydraulic actuator for extension and
retraction. A claw-like hook on the actuator serves
as a downlock for the nose gear.
5-84. OPERATION. The nose gear shock strut is
pivoted just forward of the firewall. Retraction
and extension of the nose gear is accomplished by
a double-acting hydraulic cylinder, the forward end
of which contains the nose gear downlock. Initial
action of the cylinder disengages the downlock
before retraction begins. A separate single-acting
hydraulic cylinder unlocks the nose gear uplock
hook.
TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
REMEDY
HYDRAULIC FLUID LEAKAGE FROM NOSE STRUT.
Defective strut seals and/or
defects in lower strut.
Replace defective seals; stone out
small defects in lower strut. Replace lower strut if badly scored
or damaged.
NOSE STRUT WILL NOT HOLD
AIR PRESSURE.
Defective filler valve
or valve not tight.
Check gasket and tighten loose
valve. Replace defective valve.
Defective O-ring at top of
strut.
Replace O-ring.
Result of fluid leakage at
bottom of strut.
Replace defective seals; stone out
small defects in lower strut. Replace lower strut if badly scored
or damaged.
Nose strut attachment loose.
Secure attaching parts.
Shimmy dampener lacks fluid.
Service shimmy dampener.
Defective shimmy dampener.
Repair or replace dampener.
Loose or worn steering components.
Tighten loose parts; replace
if defective.
Loose torque links.
Add shim washers and replace
parts as necessary.
Loose wheel bearings.
Replace bearings if defective;
tighten axle nut properly.
Nose wheel out of balance.
Refer to applicable paragraph.
NOSE WHEEL SHIMMY.
5-86. REMOVAL OF NOSE GEAR ASSEMBLY.
a. Jack aircraft or weight the tail of aircraft to
raise nose wheel off the ground.
~the
WARNING
Before working in landing gear wheel
wells, PULL-OFF hydraulic pump circuit breakers. Thru Serial 21062273,
the pump circuit breaker is located in
5-36
the circuit breaker panel, located immediately forward of the pilot's control
wheel. Beginning with Serial 21062274,
the pump circuit breaker is located in
circuit breaker panel, located immediately forward of the left forward
doorpost. The hydro-electric power
pack system is designed to pressurize
the landing gear DOOR CLOSE system
to 1500 psi at any time the master switch
MODEL 210 & T210 SERIES SERVICE MANUAL
I.
Work entire nose gear assembly free of aircraft.
5-87. DISASSEMBLY OF NOSE GEAR STRUT.
(Refer to figure 5-15.)
k. Remove orifice support by removing bolt at
top of strut Remove and discard O-ring from
orifice support.
1. Remove collar from upper strut. To remove
collar. remove bolt and tab washer. Remove
washers. shims. if installed, and steering collar.
NOTE
The following procedure applies to the
nose gear shock strut after it has been
removed from the aircraft, and the nose
wheel has been removed. In many cases.
separating the upper and lower struts
will permit inspection and parts INSPECTION
replacement without removal or complete
strut disassembly.
WARNING
Deflate strut completely before removing
bolt (33), lock ring (31) or bolt (2). Also
deflate strut before disconnecting torque
links.
a. (Refer to figure 5-14.) Remove torque links (17).
Note positions of washers, shims, spacers, and bushings.
b. (Refer to figure 5-14.) Remove shimmy dampener
(10) and steering bungee (12).
c. Remove link from steering shaft and collar.
d. Remove lock ring from groove inside lower
end of upper strut A small access hole is provided
at the lock ring groove to facilitate removal of lock
ring.
NOTE
Hydraulic fluid will drain from strut as
lower strut is pulled from upper strut.
a straight. sharp pull. remove lower
e. Using
strut from upper strut. Invert lower strut and
drain hydraulic fluid from strut.
f.
and bearing
f. Remove
Remove lock
lock ring
ring and
bearing from
from lower
lower
strut.
g. Slide shims, if used, packing support ring,
scraper ring, retaining ring and lock ring from
lower strut.
NOTE
Note number of shims, relative position
and top side of each ring and bearing to
aid in reassembly.
h. Remove and discard O-rings and back-up
rings from packing support ring.
i. Remove metering pin and base plug by
removing bolt from lower strut and fork assembly.
NOTE
Lower strut and fork are a press fit.
drilled on assembly. Separation of these
parts is not recommended. except for
replacement rof parts.
j. Remove and discard O-rings from metering
pIN and base plug.
5-38
Revision 3
NOTE
Upper and lower trunnions are press
fitted to the upper strut with braces
installed during assembly. Pin is also
press fitted to the lower trunnion.
AND REPAIR OF SHOCK
INSPECTION AND REPAIR OF SHOCK
5
STRUT COMPONENTS. (Refer to figure 5-15.)
a. Bushings and bearings in upper trunnion and
lower trunnion may be replaced as required.
Needle bearing in collar should not be replaced
Replace entire steering collar if needle bearing is
defective.
b. Thoroughly clean all parts in solvent and
inspect them carefully. Replace all worn or
defective parts and all O-rings, seals and back-up
rings with new parts
c. Sharp metal edges should be smoothed with
No. 400 emery paper, then cleaned with solvent.
5-89 REASSEMBLY OF NOSE GEAR STRUT.
(Refer to figure 5-15.)
NOTE
Assemble these parts lubricated with a film
of Petrolatum W-P-236, hydraulic fluid
ML-H-5606 or Dow Corning DC-7.
a. Install top washer (21), steering collar (21), shims
(22) (as many as were removed), and collar (23). Screw
collar (23) up threads on lower end of upper strut (10)
until it is flush with the lower end of the strut, to the
nearest
one-third
turn.
shimscollars.
as required
above
are
Shims
between
gap Use
to fill
lower washer,
available from Cessna Parts Distribution (CPD 2),
through Cessna Service Stations, as follows:
1243030-5
1243030-6
1243030-7
|
0.006"
.0.012"
0.020"
................ ....
............
................
NOTE
installed. secure collar (23) with bolt (43)
and secure bolt with tab washer (44) by
bending tabs of washer.
base plug (36).
b. Install O-ring (37) on base plug (36).
c. Install 0-ring (35) on metering pin (38). and
install in base plug (36).
d. Install bolt (33) through holes in fork (34) and
base plug (36). Install nut on bolt.
e. Install lock ring (31). retaining ring (30) and
scraper ring (29) down over lower strut (27).
Ensure they are installed in same positions as they
were when rmoved.
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Ensure that beveled edge of bearing is
installed up next to lock ring.
h. Install upper strut assembly over lower strut
assembly.
Install lock ring (31) in groove in lower end of
i.
upper strut (10). Position lock ring so that one of
its ends covers the small access hole in the lock
ring groove.
j. Install steering shaft (17) up through hole in
lower trunnion (8) and hole in upper trunnion (3).
k. Install steering arm (14) over steering shaft
(17) and secure with roll pins.
1. Install link (18) to bottom of steering shaft (17)
and attach opposite end to steering collar (21).
m. If braces (1) were removed, they should be
installed. connecting at upper trunnion (3) and
lower trunnion (8).
n. Attach lower torque link to torque link fitting
(32) and upper torque link to steering collar (21).
o. Install O-ring (6) and filler valve (5) on orifice
support (7).
p. Install orifice support in upper strut (10),
install bolt (2).
q. Service shock strut as outlined in Section 2 of
this manual,
5-90.
INSTALLATION OF NOSE GEAR STRUT.
WARNING
Before working in landing gear wheel
wells, PULL-OFF hydraulic pump circuit breakers. Thru Serial 21062273,
the pump circuit breaker is located in
the circuit breaker panel, located immediately forward of the pilot's control
wheel. Beginning with Serial 21062274,
the pump circuit breaker is located in
the circuit breaker panel, located immediately forward of the left forward
doorpost. The hydro-electric power
pack system is designed to pressurize
the landing gear DOOR CLOSE system
to 1500 pst at any time the mast.r switch
is turned on. Injury might occur to
someone working in wheel well area it
master switch is turned on for any
reason.
a. Work entire nose gear assembly into nose gear
wheel well.
NOTE
Trunnion bolts are accessible from inside
the cabin, at the very forward end of the
tunnel cover at the firewall. Two men
will be require to install these bolts, one
working inside the cabin. the other
working in the nose wheel well.
5-40
b. Install trunnion bolts (items 4 and 13, figure
5-14.)
c. Install nose gear strut door tie rods (items 2,
figure 5-21.) Install right-hand tie rod on outboard
side of eyebolt only (as shown in figure 5-21 ), when
connecting nose gear strut doors. Left-handtie rod
clevis should be installed as shown in figure 5-21.
d. Install nose gear actuator, washers, spring
clip and castellated nut.
NOTE
When connecting nose gear actuator to
strut, lubricate and torque bolt as outlined in the lubrication charts in Section
2 of this manual.
e. Install steering bungee to steering bellcrank.
f. Connect wires marked for identification at
safety switch on torque links, and install clamps
attaching wires to nose gear strut.
g. Connect electrical wires marked for identification at gear-down microswitch, located on forward
end of nose gear actuator (item 5, figure 5-19.)
h
Connect nose wheel door push-pull rods (items
13 figure 5-21.)
Rig nose gear and nose gear doors in accordance
1
with procedures outlined in applicable paragraphs in
step cable in accordance with
retractable
J. Rig retractable step cable in accordance with
procedures outlined in applicable paragraph in this
Section.
5-91.
16.)
SHIMMY DAMPENER.
(Refer to figure 5-
5-92. DESCRIPTION. The shimmy dampener is a
self-contained hydraulic cylinder which acts as a
restrictor. When the steering system reacts too
rapidly, the shimmy dampener maintains pressure
against the steering arm by means of a piston
which permits a restricted flow of hydraulic fluid
from either end of the cylinder to the other through
an orifice in the piston.
5-93. SHIMMY DAMPENER REMOVAL (Refer to
figure 5-14.)
a. Remove bolt securing shimmy dampener to
steering shaft.
b. Remove bolt attaching dampener to bracket.
attached to lower trunnion.
c. Remove shimmy dampener from aircraft.
5-94. DISASSEMBLY OF SHIMMY DAMPENER.
(Refer to figure 5-16.)
a. Remove outer retaining ring (7).
b. Remove bearing head (6).
c. Remove O-rings (3) from bearing head.
d. Remove internal retaining ring (5).
e. Remove rod assembly (8).
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
THIS INSTALLATION LOCATED AT
EXTREME TOP FORWARD OF NOSE
GEAR WHEEL WELL.
2
2.
3.
4.
5.
6.
RH Tunnel Wall
Bellcrank and Hook Assembly
Bracket (on opposite side of hook)
Uplock Switch
Inner Bearing Race
8.
9.
10.
11.
Figure 5-18.
Bearing
Spring
Actuator
Links
Nose Gear Uplock Mechanism
5-102. NOSE GEAR UPLOCK MECHANISM.
(Refer to figure 5-18.)
5-103. DESCRIPTION. The nose gear uplock
mechanism, located in the top of the nose wheel
well, is a hydraulically-unlocked hook that is
spring-loaded to the locked position. The nose
gear indicator switch is attached to a bracket
welded to the uplock hook.
5-104. REMOVAL OF NOSE GEAR UPLOCK
MECHANISM. (Refer to figure 5-18.)
a. With master switch OFF, pump landing gear
doors open.
NOTE
With doors open, all components are
readily accessible at top forward end of
the nose wheel well.
b. Disconnect links (11) from actuator (10).
c. Disconnect spring (9) from aircraft structure
or from hook on bellcrank assembly (3).
d. Unscrew nut attaching uplock switch (5).
e. Remove bolt (1) through right-hand tunnel
wall.
position and install washer between bellcrank and
right-hand tunnel wall, then install bellcrank and
hook assembly; install bolt (1), bearing (8), washer
and nut.
b. Install uplock switch (5).
c. Attach spring (9) to aircraft structure or to
hook on bellcrank assembly (3).
d. Connect links (11) to actuator (10).
e. Rig system in accordance with applicable
paragraph.
5-106. NOSE GEAR DOWNLOCK MECHANISM.
(Refer to figure 5-19.)
5-107. DESCRIPTION. The nose gear downlock
mechanism is a hook at the piston rod end of the
nose gear actuator.
5-108. REMOVAL AND INSTALLATION OF
NOSE GEAR DOWNLOCK MECHANISM. (Refer
to figure 5-19.) Refer to figure 5-20 and paragraph
5-111, which outlines procedures for removing the
nose gear actuator. Components of the downlock
mechanism will be freed as the actuator is
removed.
5-109.
NOSE GEAR ACTUATOR.
(Refer to figure
5-20.)
5-105. INSTALLATION OF NOSE GEAR UPLOCK
MECHANISM. (Refer to figure 5-18.)
a. Place bellcrank and hook (3) assembly in
5-110. DESCRIPTION. The nose gear actuator
extends and retracts the nose gear and serves as a
5-43
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
*May be purchased from:
Electro-Film Inc.
7116 Laurel Canyon Blvd.
Hollywood, CA 91605
* May be purchased from:
Everlube Corp.
P.O. Box 2200
Hi-Way 52 N.W.
West LaFayette, Ind. 47906
The downlock hooks (2) and (7) have been dry
film lubricated at the factory and should last
the life of the parts. However, they may be
field lubricated with the following products:
* 1. Lubri-Bond A.
2. Lubri-Bond 220.
* 3. Permasilk.
After application allow parts to air dry for
six hours, or dry for one hour at 120°F.
1.
2.
3.
4.
Lower Trunnion
Hook
Crossbar
Rod End
6. Actuator
7. Hook
8. Bolt
9. Shimmy Dampener
/>
Figure 5-19.
5-44
Revision 2
Nose Gear Downlock Mechanism
MODEL 210 & T210 SERIES SERVICE MANUAL
inches under a 19.80 + 2.0 pound load.
c. Inspect hooks, spring guide. bearing end.
piston, cylinder and bushing for cracks, chips,
scratches, scoring, wear or surface irregularities
which may affect their function or the overall
function of the nose gear actuator.
d. Repair of most parts of the actuator assembly
is impractical. Replace defective parts with
serviceable parts.
e. Minor scratches and scores may be removed
by polishing with fine abrasive crocus cloth
(Federal Specification PC-458), providing their
removal does not affect operation of the unit.
b. Disconnect and cap or plug hydraulic lines at
actuator.
c. Disconnect and tag up-limit switch electrical
wires.
d. Remove cotter pin and clevis pin attaching
actuator link to bellcrank arm. Note position of
spacer washers and direction of clevis pin.
e. Remove nuts, washers and bolts attaching
actuator to wheel well tunnel wall. Note and retain
shims between actuator and tunnel wall
f. Remove bolt, washer and nut attaching
bellcrank at top of nose wheel.
NOTE
5-114. ASSEMBLY OF NOSE GEAR ACTUATOR.
(Refer to figure 5-20.)
NOTE
When reassembling actuator, install new
Wrings
0-rings and
and back-up
back-up rings
rings lubricated
lubricated with
with
a film of Petrolatum VV-P-236, hydraulic
fluid MIL-H-5606, or Dow-Corning DC-7.
a.
Install O-rings and back-up rings in bearing
end.
end.
b. Install O-rings and back-up rings on piston.
c. Insert piston into cylinder. Do not damage
back-up rings and O-rings when inserting piston.
d. With knurled nut on cylinder, install bearing
end on cylinder. Use care to avoid damage to Orings and back-up rings when installing bearing
end on cylinder.
NOTE
Centerlines of hook pin holes and bushing hole
must be parallel within . 005 with actuator
assembled to a length of 11. 58" ±. 03 (thru
1978 models.) 11. 98" *. 03 (Beginning with
1979 models).
e. Tighten and safety wire knurled nut.
f. Install lock nut on end of piston.
g. Assemble and install hook assembly on
piston.
5-115. INSTALLATION OF NOSE GEAR
ACTUATOR.
NOTE
Before installing nose gear actuator,
check condition of fit and attaching bolts
and bushings. Replace any defective
parts. Fill actuator with hydraulic fluid.
a. Attach aft end of actuator to fuselage structure
with bolt, washer and nut. Safety nut with cotter
pin.
b. Assemble and attach nose gear downlock
mechanism to lower trunnion as shown in figure 518.
5-116. REMOVAL AND INSTALLATION OF
NOSE GEAR UPLOCK AND RELEASE
ACTUATOR.
a. Disconnect uplock spring.
5-46
Revision 2
Use care to avoid dropping bearings in
bellcrank assembly. Retain washers used
as shims at each end of bellcrank.
g. Install uplock mechanism and actuator by
reversing the preceding steps. Install shims and
washers as noted during removal.
REPAIR OF PARTS AND REASSEMBLY OF
REPAIR~ OF PARTS AND REASSEMBLY OF
NOSE GEAR UPLOCK AND RELEASE
ACTUATOR.
Refer to figure 5-5 and paragraphs
ACTUATOR. Refer to figure 5-5 and paragraphs
5-118. NOSE GEAR DOOR SYSTEM.
figure 5-21 )
(Refer to
5-119. DESCRIPTION. The nose gear door system
consists of a right and left forward door, actuated
by push-pull rods and a torque tube assembly and a
right and left aft door, mechanically linked to the
nose gear trunnion.
5-120. OPERATION. The nose gear forward doors
open for nose gear retraction or extension and
close again when the cycle is completed. These
doors are held in the closed position by the door
lock valve, located in the door manifold assembly,
mounted on the power pack, by trapping fluid in
the door lines. Actuation of the nose gear forward
doors is accomplished by a double-acting hydraulic
cylinder. The nose gear aft doors are mechanically
linked to the nose gear trunnion. these doors open
as the nose gear extends and close as it is
retracted.
5-121. REMOVAL AND INSTALLATION OF
NOSE WHEEL DOORS. (Refer to figure 5-21.)
a. Open landing gear doors.
b. Remove engine cowl.
c. Disconnect push-pull rod from bracket on door
by removing nut, bolt and washers.
d. Remove nuts and bolts attaching each hinge
pivot. Work from upper side of cowl opening to
remove bolts. Retain bushings in hinge pivot.
e. To replace nose wheel doors, reverse the
preceding steps.
5-122. REMOVAL AND INSTALLATION OF
NOSE WHEEL DOOR MECHANISM. (Refer to
figure 5-21. )
a. Open landing gear doors.
b. Disconnect actuator at torque tube by
MODEL 210 & T210 SERIES SERVICE MANUAL
1 2
CLEVELAND NOSE WHEEL
NOTE
Tighten nuts (12) evenly and
torque to 90 lb in.
6
12
NOTE
Tighten nuts (16) evenly and
torque to 140-150 lb in.
2
Do not use impact wrenches
on thru-bolts (20) or nuts (16).
13
15
14
17
9
McCAULEY NOSE WHEEL
1. Snap Ring
2. Grease Seal Ring
3. Bearing
5.
6.
7.
8.
9.
10.
Tube
Grease Seal Felt
Thru-Bolt
Bearing Cup
Male Wheel Half
Female Wheel Half
11.Washer
12.
Nut
13. Retainer Ring
14. Grease Seal Retainer
15. Felt Grease Seal
16. Nut
17. Washer
14
18. Wheel Half
19. Bearing Cup
20. Thru-Bolt
21. Bearing Cone
22. Tube
23. Tire
Figure 5-23.
15
23
Nose gear Wheel and Tire Assembly
NOTE
Use of recapped tires or new tires not listed
on the aircraft equipment list are not recommended due to possible interference between
the tire and structure when landing gear is in
the retracted position.
5-131. OPERATION. The nose gear wheel is freerolling on an independent axle and is used to steer
the aircraft while taxiing by means of the nose
wheel steering system.
5-132. REMOVAL OF NOSE WHEEL AND TIRE
ASSEMBLY.
5-49
MODEL 210 & T210 SERIES SERVICE MANUAL
a. Weight tail of aircraft to raise nose wheel off
the ground.
b. Remove nose wheel axle bolt.
c. Use a rod or long punch inserted in ferrule to
tap opposite ferrule out of nose wheel fork.
d. Remove spacers. axle tube and hub caps
before disassembling nose wheel.
e. Reverse preceding steps to install nose
wheel. Tighten axle bolt until a slight bearing
drag is obvious when the wheel is turned. Back off
nut to nearest castellation and install cotter pin.
5-133. DISASSEMBLY OF CLEVELAND NOSE
WHEEL AND TIRE ASSEMBLY. (Refer to figure
5-23.)
WARNING
Injury can result from attempting to
separate wheel halves with te tire
inflated. Avoid damaging whe
when breaking tire beads loose.
a.
and
b.
c.
d.
seal
Remove valve core. completely deflate tire.
break tire beads loose.
Remove thru-bolts and separate wheel halves.
Remove tire and tube.
Remove snap rings (1), grease seal felts (6), grease
rings (2), and bearings (3).
~~~~~~NOTE
~not
The bearing cups are a press fit in the wheel
halves and should not be removed unless replacement is necessary. To remove, heat
wheel half in boiling water for 15 minutes.
Using an arbor press, if available, press
out bearing cup and press in the new one
while the wheel is still hot.
f. Inflate tire to seat tire beads, then adjust to
correct pressure.
5-136. DISASSEMBLY OF McCAULEY NOSE WHEEL
AND TIRE ASSEMBLY. (Refer to fieure 5-23.)
a. Remove hub caps, completely deflate tire, and
break tire beads loose at wheel flanges.
WARNING
Injury can result from attempting t
remove
wheel flanges with tire and tube inflated.
Avoid damaging wheel flanges when breaking
tire beads loose. A scratch, gouge or nick
in wheel flange could cause wheel failure.
b. Remove nuts and washers.
c. Remove thru-bolts and washers.
d. Separate and remove wheel halves from tire and
tube
e. Remove retainer ring (13), grease seal retainer (14),
elt grease seal (15), and bearing cone (21) from each
wheel half 18)
wheel half(18).
NOTE
The bearing cups (races) are a press fit in
the wheel hub and should not be removed
unless a new part is to be installed. To
remove the bearing cup, heat wheel hub in
boiling water for 30 minutes, or in an oven
to exceed 121°C (250°F). Using an
arbor press, if available, press out the
bearing cup and press in the new bearing
cup while the wheel hub is still hot.
5-137. INSPECTION AND REPAIR OF McCAULEY
NOSE WHEEL AND TIRE ASSEMBLY.
a. Clean all metal parts and felt grease seals in
Stoddard solvent, or equivalent, and dry thoroughly.
NOTE
5-134. INSPECTION AND REPAIR OF CLEVELAND
NOSE WHEEL AND TIRE ASSEMBLY. Procedures
outlined in paragraph 5-41 for the main wheel and
tire assemblies may be used as a guide for inspection and repair of the nose wheel and tire assembly.
5-135. REASSEMBLY OF CLEVELAND NOSE
WHEEL AND TIRE ASSEMBLY. (Refer to figure
5-23.)
a. Place tube inside tire and align balance marks
on tire and tube.
b. Place tire and tube on wheel half with tube valve
stem through hole in wheel half.
CAUTION
Uneven or improper torque of the thru-bolt
nuts may cause bolt failure with resultant
wheel failure.
c. Insert thru-bolts, position other wheelhalf
and secure with nuts and washers. Torque bolts to
value stipulated in figure 5-23.
d. Clean and repack bearing cones with clean wheel
~~~~~bearing
grease.
e. Assemble bearings (3), grease seal rings (2), and felt
grease seal felts (6) into wheel halves and install snap
rings (1).
5-50 5-0
Revision
Revision 3
A soft bristle brush may be used to remove
hardened grease. dust or dirt
b. Inspect wheel halves (18) for cracks or damage.
c. Inspect bearing cones (21), bearing cups (19),
retainer rings (13), and felt grease seals (15) for wear or
damage.
d. Inspect thru-bolts (20) and nuts (16) for cracks in
threads or cracks in radius under bolt head.
e. Replace cracked or damaged wheel halves (18).
f. Replace damaged retainer rings (13)and seals.
g. Replace any worn or cracked thru-bolts (20) or nuts
(16).
h. Replace any worn or damaged bearing cups (19) or
bearing cones (21).
i. Remove any corrosion or small nicks.
j. Repairreworked areasof wheel bycleaning
thoroughly, then applying one coat of clear lacquer paint.
Section 2 ofthis manual.
5-138. REASSEMBLY OF McCAULEY NOSE WHEEL
TIRE ASSEMBLY. (Refer to figure
5-23.
a. bearing
Assemble
cone. grease seal retainer.
seal
grease seal
retainer and retainer
ring into both wheel halves.
b. Insert tube in tire, aligning index marks on tire
and tube.
MODEL 210 &T210 SERIES SERVICE MANUAL
c. Place wheel half into tire and tube (side opposite
valve stem), aligning base of valve stem in valve slot.
With washer under head of thru-bolt, insert bolt
through wheel half.
d. Place wheel half into other side of tire and tube.
aligning valve stem in valve slot.
e. Install washers and nuts on thru-bolts and pretorque to 10-50 b. in.
CAUTION
5-158. HAND TOOLS. The following hand tools
are necessary for repair work on the power pack
and other hydraulic components.
Snap Ring Pliers
Strap Wrench (for removing door solenoids and
various cylinder barrels of the hydraulic
actuators.)
Needle-Nose Pliers
Pin Punches
Duck-bill Pliers
Box end and Open end Wrenches
Locality-faricated item handy for power pack
Uneven or improper torque of nuts can cause
Locality-fabricated items. handy
failure of bolts with resultant wheel failure.
aluminum rods. ground
repair. are various 1/4inch
Do not use impact wrench on thru-bolts or nuts.
to a gradual taper. and hooks formed from brass
from
welding rod to extricate small plungers
Hooks from brass welding
welding rod to extricateformed
f. Prior to torquing nuts, inflate tire to 10-15 psi
hydraulic ports. Hooks formed from brasssowelding
air pressure to seat tire,
as not
rod must not be over 1/16-inch in length,
Dry torque nuts evenly to 140-150 in lb.
to scratch or score the bore. Various sizes of Alien
g. Dry torque nuts evenly to 140-150 in lb.
wrenches may be welded to '"T handles for use
h. Inflate tire to pressure specified in Section 1.
when removing, installing or adjusting the various
internal wrenches. plugs or valves.
ASSEMBLY.
TIRE
5-139
TIRE ASSEMBLY.
5-159. COMPRESSED AIR The simplest method
a. Install nose wheel in fork and install ferrules.
of removing some hydraulic parts in inaccessible
b. Install axle stud.
galleries of the power pack is a quick blast of
c. Tighten axle stud until a slight bearing drag is
compressed air from behind. Parts can be blown
obvious when the wheel is turned. Back off nut to
out in seconds, which would otherwise take endless
nearest castellation and install cotter pins.
"fishing" operation to extricate. An air hose and
nozzle are common-sense tools.
5-140. THRU 5-151. DELETED,
5-152.
HYDRAULIC POWER SYSTEM
LEAK CHECK. (Refer
COMPONENTS. (Refer to figure 5-24.)
5-159A. HYDRAULIC SYSTEM LEAK CHECK (Refer
to figure 5-24.)
5-153. GENERAL DESCRIPTON. The hydraulic
power system includes equipment required to
provide a flow of pressurized hydraulic fluid to the
retractable landing gear system. Main components,
of the hydraulic power system include the power
pack and the emergency hand pump.
a. Jack aircraft in accordance with procedures in
Section 2 of this manual.
b. To relieve system pressure, pull the GEAR PUMP
circuit breaker to OFF, move the gear selector handle to
UP, and move back to the DOWN position.
c. Install a 0-2000 PSI gage at the tee (Index 47, figure
5-26) on the left side of the power pack.
d. Push the GEAR PUMP circuit breaker to the ON
position, turn ON the master switch, and move gear
selector handle to the UP position.
e. Monitor pressure gage, after retraction cycle is
complete, for pressure bleed down.
f. If bleed down occurs, it can be an internal or
5-154. HYDRAULIC COMPONENTS REPAIR
Since emphasis here is on repair and not overhaul
of the basic components of the hydraulic system. it
is unlikely that the mechanic will go through all of
the procedures outlined. Instead. he will repair the
particular item which is causing the difficulty.
5-155. REPAIR VERSUS REPLACEMENT. Often.
the moderate trade-in price for a factory-rebuilt
component is less than the accumulated cost of
labor, parts and (often time consuming) trial and
error adjustment. Repair or replacement of a
component will depend on the time, equipment and
skilled labor that is locally available.
external leak anywhere in the system.
NOTE
When any line is disconnected, be prepared
for fluid leakage.
g. Disconnect the return line from the gear selector. If
fluid comes from the selector, the internal leak is in the
5-156. REPAIR PARTS AND EQUIPMENT. Repair
parts may be ordered from the applicable Parts Catalog.
system.
h. If no leak-by is found, it can be assumed there is an
available from Cessna Parts Distribution (CPD 2)
through Cessna Service Stations.
i. Power pack internal leakage can only be attributed
to a bad thermal relief valve, self-relieving check valve,
or self-relieving check valve O-ring. The only way to
Test equipment may be ordered from the Special Tools
and Support Equipment Catalog. Both publications are
5-157.
EQUIPMENT AND TOOLS.
internal leak in the power pack. If leak is found, proceed
to step "j." Reconnect the return line.
isolate part that is leaking is to systematically replace
Revision 3
5-51
MODEL 210 & T210 SERIES SERVICE MANUAL
the self-relieving check valve O-ring, self-relieving check
valve, and then thermal relief valve. Repeat leak test
after replacement of each part to ensure leak correction.
j. Remove gear DOWN line from selector. If fluid
comes from the line, one or more of the gear actuators is
leaking. To locate the leaking actuator, disconnect the
return line from each actuator; the leaking actuator will
have fluid draining from the actuator port. Following the
appropriate paragraphs in this section, remove, overhaul,
and reinstall the actuator.
k. Reconnect gear DOWN line to the selector.
1. Recheck all lines that were disconnected for
security.
m. Lower the landing gear. Following the procedures
in step "b.", relieve the system pressure.
n. Remove the pressure gage from service tee.
o. In accordance with the procedures in Section 2 of
this manual replenish the power pack reservoir with
MIL-H-5606 hydraulic fluid and bleed the system.
p. Remove aircraft from jacks.
5-160.
POWER PACK.
5-161. DESCRIPTION. The hydraulic power pack,
located in the pedestal, is a multi-purpose control unit. It
contains a hydraulic reservoir, valves, an electricallydriven motor, and the pump. An emergency hand pump,
located between the pilot's and copilot's seats, uses
reservoir fluid to permit manual extension of the landing
gear.
NOTE
The hydraulic power pack relief valve,
thermal relief valve, and pressure switch can
be operationally checked on the aircraft
without power pack removal from the aircraft
or disassembly. Refer to paragraph 5-161A
for specific instructions. Refer to paragraph
5-172A for relief valve and thermal relief
valve bench check instructions if the power
pack is removed from aircraft.
(6) Push landing gear circuit breaker in; power pack
should run; monitor pressure.
(7) Relief valve should open at 1800 PSI, + 0 or -50
PSI.
(8) After check is complete, remove pressure from
system, remove pressure gage, install cap on tee (47),
pull landing gear circuit breaker, remove jumper wire,
push landing gear circuit breaker back in, and return
system to original configuration.
b. Thermal Relief Valve.
(1) With aircraft onjacks and pressure gage
installed at tee (47) fitting on left side of power pack,
pull landing gear circuit breaker.
(2) Select landing gear to DOWN position.
(3) Extend emergency gear pump handle.
(4) Pump emergency gear pump handle and monitor
pressure. Thermal relief valve should open at 2050
PSI ± 100 PSI.
(5) After check is complete, remove pressure from
system, remove pressure gage, and install cap on tee
(47).
(6) Push in landing gear circuit breaker, and return
system to original configuration.
c. Pressure Switch.
(1) With aircraft on jacks and pressure gage
installed at tee (47) fitting on left side of power pack,
pull landing gear circuit breaker.
(2) Select landing gear UP and DOWN several times
to relieve pressure in landing gear system.
(3) Select landing gear UP, and push in landing gear
circuit breaker.
(4) After gear raising cycle is complete, check
pressure. Pressure should be 1500 PSL
(5) Select gear DOWN. After gear lowering cycle is
complete, pressure should be 1500 PSI.
(6) After check is complete, remove pressure from
system, remove pressure gage, install cap on tee, and
return system to original configuration.
5-162.
5-25.)
REMOVAL OF POWER PACK. (Refer to figure
NOTE
5-161A. ON-AIRCRAFT HYDRAULIC POWER PACK
OPERATIONAL CHECKS. (Refer to figure 5-26.)
The relief valve, thermal relief valve, and pressure
switch should be pressure checked each 100 hours. They
can be operationally checked without removal from
aircraft. For bench check instructions after removal from
power pack, refer to paragraph 5-172A.
NOTE
Checks are to be performed with external
power set at 28.5 volts.
a.
Relief Valve.
(1) Jack aircraft in accordance with procedures
outlined in Section 2.
(2) Remove cap and install pressure gage at tee (47)
fitting on left side of power pack.
(3) Pull landing gear circuit breaker.
(4) Select landing gear handle to DOWN position.
(5) Install 18 gage (minimum)jumper wire between
buss side of contactor and small terminal on pump
motor contactor (to energize coil).
5-52
Revision 3
As hydraulic lines are connected or removed,
plug or cap all openings to prevent entry of
foreign material in the lines or fittings.
a. Remove front seats and spread drip cloth over
carpet.
b. Remove decorative cover from pedestal as outlined
in Section 9 of this manual.
c. Remove upper panel from aft face of pedestal panel.
d. Remove screws attaching indicator assembly at top
of pedestal; remove indicator assembly.
e. Remove four bolts attaching wheel and gear box
assembly; remove wheel and gear box assembly.
f. Loosen idler sprocket by loosening bolt and sliding
sprocket inboard in slot.
g. Disconnect chain at its connecting link.
h. Remove left-hand and right-hand chain guards.
i. Allow chain to remain on gimbal assembly in lower
pedestal area.
j. Position gallon container under drain elbow at righthand forward side of pedestal.
k. Remove cap from elbow and attach drain hose.
MODEL 210 & T210 SERIES SERVICE MANUAL
MAIN GEAR
DOOR ACTUATOR
ACCUMULATOR
MAIN GEAR
ACTUATOR
UNLOCKACTUATOR
SELECTOR VALVE
POWER PACK
DOOR ACTUATOR
X
ACTUATOR
AGEAR
NOSE
DOOR ACTUATOR
21062274 thru 21062954
. Figure 5-24.
Hydraulic Syatem Components (Sheet 2 of 3)
5-53
MODEL 210 &T210 SERIES SERVICE MANUAL
5-163. DISASSEMBLY OF POWER PACK. (Refer to
figure 5-26.)
a. Remove fittings from body assembly and place body
assembly in vise.
b. Remove nut (23), reservoir washer (22), and packing
(3) at stud (31) at bottom of reservoir (25); remove
reservoir.
NOTE
If reservoir will not disengage from body
assembly, replace fittings and cap or plug all
fittings except vent fitting. Attach air hose at
vent fitting and apply pressure (not to exceed
15 PSI: reservoir proof pressure); remove
reservoir. A strap clamp is not recommended
as clamp may damage reservoir.
c. Remove door manifold assembly (Index 35, figure
5-27) and gear solenoid assembly from body assembly of
power pack.
NOTE
Disassembly of pressure switch assembly and
relieve valve assembly is normally not
required. Refer to applicable paragraphs for
specific instructions.
d. Remove pressure switch and dipstick from body
assembly.
e. Remove large packing (3) from bottom of body
assembly.
f. Remove baffle (29), spacers (27), and washer (26).
g. Remove union (14), packing (3), retainer ring(7),
and screw (24) at bottom of reservoir (25).
h. Remove motor and pump assembly (10) from body
assembly.
i. Remove packings and back-up rings from pump
assembly (10); remove coupling (11).
j. Remove return tubes (30) and packings from body
assembly.
k. Remove relief valve assembly from body assembly.
m. Remove fittings from body assembly, if still
installed, union (14), packing (3), retainer ring (7), and
fluid filter screen (8) from body assembly.
n. Remove thermal relief valve and check selfrelieving check valve from body assembly.
NOTE
To remove thermal relief valve when
power pack is installed in aircraft,
remove retainer (6). While holding your
hand to catch valve, gently pump hand
pump. Valve will be ejected out into your
hand. Be careful not to pump hand pump
too hard.
5-164. INSPECTION AND REPAIR OF POWER
PACK COMPONENTS.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661, or equivalent) and dry with
filtered air.
b.
Inspect seating surfaes. They should have
very sharp edges. Seats may be lapped, if
necessary, to obtain sharp edges.
c. Inspect all threaded surfaces for serviceable
condition and cleanliness.
d. Inspect all parts for scratches, scores, chips,
cracks and Indications of excessive wear.
5-165. REASSEMBLY OF POWER PACK. (Refer
to figure 5-26.)
NOTE
Lubricate threads, new packings and
retaining rings with a film ofPetrolatum
VV-P-236,
hydraulic fluid MILH-5606, or
Dow-Corning DC-7 during
reassemblyof power pack
a Assemble and install thermal relief valve and
self relieving check valve in body assembly.
c.
Install fluid filter screen (8), retainer ring (7),
packing (3) and union (14) in top of body assembly (34).
c. Install suction screen assembly (32), if removed.
NOTE
S uction screen assembly (32) need not be
removed from body assembly to be cleaned
However, if suction screen assembly is
damaged, it should be removed as outlined in
step "1." of this paragraph observing the
following caution:
CAUTION
Use extreme caution in removing suction
screen assembly. Damage to suction screen
assembly or clearance between suction screen
assembly and body assembly will cause slow
landing gear retraction.
l. Working through center hole in top of body
assembly, and using a drift or punch made of soft
material, tap out suction screen assembly (32).
5-56
Revision 3
CAUTION
Use extreme caution when installing
suction screen assembly.
Damage to
screen assembly or clearance between
screen assembly and body will cause slow
landing gear retraction.
d. Install relief valve assembly in body
assembly.
e. Install packings and return tubes (30) in body
assembly.
f. Install packings and back-up rings on pump
assembly (10); install coupling (11).
g. Install pump assembly (10) and motor on body
assembly.
h. Install screen (24), retainer ring (7), packing (3),
and union (14) on bottom of reservoir (25).
i. Install washer (26), spacers (27), and baffle (29).
j. Install large packing (3) on bottom of body assembly.
MODEL 210 & T210 SERIES SERVICE MANUAL
SHIM (39) APPLICABILITY
SHIM PART NO.
9880705-1
9880705-2
9880705-3
THICKNESS
.005
.010
.016
EFFECT IN
MATERIAL PRESSURE (PSI)
BRASS
BRASS
BRASS
60
120
200
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
k. Install dipstick (9), pressure switch, door manifold
assembly (Index 35, figure 5-27), and gear manifold
assembly on body assembly.
1. Attach reservoir (25) to body assembly with packing
I (3), reservoir washer (22), and nut (23).
The chart in figure 5-26 lists shims (39)
by part number, thickness and effect on
operating pressure (psi).
5-166. INSTALLATION OF POWER PACK. (Refer
a. Work power pack into position and install
three bolts that secure power pack to pedestal.
b. Connect all hydraulic lines to power pack
fittings. Ensure that all fittings are properly
installed, with jamnuts tight, after lines are
tightened.
Do not damagethreads of fitting (45) are
of fitting (45)primer
and
Loctite Grade
primed with
sealed with Loctite Grade Av sealer and
sealed with Loctite Grade AV sealer.
(Refer to figure 5-
f. Remove piston (43).
g. Remove packings (42) and (44).
h. Remove snubber (46) from fitting (45).
CAUTION
Threads of snubber (46) and fitting (45)
are primed with Loctite Grade T primer
sealed with Loctite Grade AV sealer.
5-170. CLEANING, INSPECTION AND REPAIR
OF PRESSURE SWITCH. (Refer to figure 5-26.)
a. Clean sealant from threads of snubber (46),
fitting (45) and guide (41) with wire brush.
b. Clean all parts with cleaning solvent (Federal
Specification P-S661, or equivalent) and dry
thoroughly.
c. Discard all removed packings (42) and (44) and
replace with new packings.
d Inspect all pressure switch parts for scratches.
scores, chips, cracks and indications of wear.
e. All damaged parts shall be replaced with new
parts.
NOTE
5-168. DESCRIPTION. When installed in the
aircraft, the pressure switch is mounted on the
right-hand (aft) side of the power pack in the
console. This switch senses pressure in the
DOOR-CLOSE line. After gear extension or
retraction (after the doors close), pressure builds in
the DOOR-CLOSE line. At approximately 1500 PSI,
the pressure
Atturning
approximately
1500 PSI,
the
pressure switchopen.
opens,
off the power
pressure switch
will continue tohold
pack. The
the electrical circuit open until pressure in the
system drops to a preset value, at which time, the
pump will again operate to build up pressure to171
approximately 1500 PSI..
NOTE
The hydraulic power pack relief valve,
thermal relief valve, and pressure switch can
each be operationally checked on the aircraftslotted
without disassembly. Refer to paragraph
5-161A for specific instructions.
5-169. DISASSEMBLY OF PRESSURE SWITCH.
(Refer to figure 5-26.)
a. Remove pin (37).
a. Remove pin (37).
b. Unscrew cap and housing assembly (36) from
fitting
fitting (45).
(45).
c. Remove spring (38).
d. Remove shims (39) from flange of guide (41).
5-58
Revision 3
Unscrew guide (41) from fitting (45).
CAUTION
c. Install wheel and gear box assembly and
indicator assembly in top of pedestal.
d. Install left-hand and right-hand chain guards
for rudder trim chain.
e. Connect chain at connecting link after
stringing chain over idler sprocket.
f. Tighten idler sprocket by sliding sprocket outboard
in slot and tightening bolt.
g. Connect ground wire to pressure switch and wire to and
motor.
h. Connected power pack wiring to plug.
i. Install upper panel on pedestal.
j. Fill reservoir on right-hand side of power
pack with clean hydraulic fluid in accordance with
procedures outlined in Section 2 of this manual.
k. Jack aircraft as outlined in Section 2 of this
manual
manual..
1. Operate gear thru several cycles to bleed system
Check for correct operation and signs of fluid leakage.
A 28V power supply should be used to augment the
ship's battery.
5-167. PRESSURE SWITCH.
26.)
e.
Thorough cleaning is important
Dirt
and chips are the greatest single cause of
malfunctions in hydraulic systems.
Carefulness and proper handling of parts
to prevent damage must be observed at
all times.
f. Snubber (46) can be cleaned with solvent, then
blown out with high pressure compressed air.
g. Assure that .062-inch vent hole is open in stop
(40)
ASSEMBLY OF PRESSURE SWITCH.
5-171. ASSEMBLY OF PRESSURE SWITCH.
(-Refer to figure 5-26.)
a. Prime threads of snubber (46) and internal
threads of fitting (45) with Loctite Grade T primer
and apply Loctite Grade AV sealer to threads of
snubber (46). Install snubber into fitting with a
screwdriver.
NOTE
When reassembling pressure switch, install
new packing and internal parts, except as
noted, lubricated with a film of Petrolatum
W-P-236, hydraulic fluid MIL-H-5606, or
-P236, hydraulic fluid MIL-5606 or
Dow-Corning DC-7.
b.
Install packing (42) in fitting (45).
MODEL 210 & T210 SERIES SERVICE MANUAL
c. Lubricate packing (44) and guide (41) and install
packing on guide.
d. Prime threads of guide (41) and internal threads
of fitting (45) with Loctite Grade T primer and apply
Loctite Grade AV sealer to threads of guide (41).
Install guide into fitting and finger tighten.
e. Install test gage in power pack body fitting.
f. Assure that sealant in fitting (45) is dry; screw
fitting assembly in console.
g. Pump emergency hand pump just enough for fluid
to seep from top of guide (41).
h. Lubricate piston (43) and insert piston into hole
in guide (41).
i. Lubricate stop (40) and install over guide (41).
j. Install exact number and thickness of shims (39)
as were removed.
NOTE
NOTE
~thermal
If same number of shims (39) are installed
as were removed, pressure should not require adjustment. If readjustment is necessary, a chart of shim part numbers, thickness and effect in pressure adjustment is
illustrated in figure 5-26.
i. If switch opens electrical circuit to solenoid at higher
than 1500 ± 50 PSI, disassemble pressure switch down to
shims (39), and remove shims as necessary to obtain
desired pressure; repeat steps "b." and "c".
j. Turn off master switch.
k. Drive new pin (37) through slot in housing skirt and
hole in fitting (45).
1. Remove aircraft from jacks.
5-172. RELIEF VALVE AND THERMAL RELIEF
VALVE ASSEMBLIES. (Refer to figure 5-26.) The
relief valve assembly (5) serves to limit that amount of
pressure which can be generated by the pump assembly
(10). The thermal relief valve (2), located on the system
side of the self-relieving check valve (1), serves to limit
the system pressure. System pressure can increase due to
expansion.
5A-172A. BENCH CHECK OF RELIEF VALVE AND
THERMAL RELIEF VALVE. (Refer to figure 5-26.)
NOTE
The hydraulic power pack reliefvalve,
thermal relief valve, and pressure switch can
be operationally checked on the aircraft
without power pack removal from the aircraft
or disassembly. Refer to paragraph 5-161A
for specific instructions.
k. Lubricate spring (38) and install over shims (39)
1. Screw cap and housing assembly (36) on fiting
(45).
NOTE
Do not install pin (37) until pressure
adjustment has been checked.
5-172. ADJUSTMENT OF PRESSURE SWITCH.
(Refer to figure 5-26.)
a. Jack aircraft in accordance with procedures
outlined in Section 2 of this manual.
b. Screw cap and housing assembly (36) on
fitting (45) enough to bottom piston (43) out in stop
(40).
c. Turn cap and housing assembly (36) back from full
thread engagement one turn, plus 0, minus one-fourth
turn, to locate hole in fitting (45) in slot in skirt of cap
and housing assembly.
d. Attach electrical connections to pressure switch,
and attach external power source.
e. Turn on master switch.
f. Pump hand pump to obtain 1500 PSI on test gage.
g. The switch should open the electrical circuit to the
pump solenoid when pressure in the system increases to
I approximately 1500 PSI.
h. If switch opens electrical circuit to solenoid
prematurely, disassemble pressure switch down to
shims (39) and add shims as necessary to obtain
desired pressure: repeat steps "b" and "c".
NOTE
The chart in figure 5-26 lists shims by
part number, thickness and the effect in
psi each shim will have on switch
operation.
If on-aircraft pressure checking of the power pack
reveals out-of-tolerance relief valve opening, it may be
necessary to determine if relief valve disassembly or
adjustment is necessary. Once removed from power pack,
individual relief valves can be bench checked.
NOTE
Adequate precautions should be taken to
recover hydraulic fluid which will be expelled
from the primary relief valve while under
pressure.
a.
Relief Valve.
(1) Using a hydraulic pump with a flow rate of 0.5 to
0.7 gallons per minute connected to a hydraulic
reservoir, a pressure gage with 2500 PSI capacity, and
a hose with appropriate fittings, connect hydraulic
pump to adapter (15) of the relief valve.
(2) Apply pressure slowly to ensure that relief valve
assembly opens at correct pressure reading. Relief
valve should open at 1800 PSI, + 0 or -50 PSI. Refer to
paragraph 5-172D for adjustment instructions.
b. Thermal ReliefValve.
1 ) Using a hand pump connected to a hydraulic
reservoir, a pressure gage with 2500 PSI capacity, and
a hose with appropriate fittings, connect hand pump to
adapter (2) of the thermal relief valve.
(2) Manually pump pressure up slowly to ensure
that relief valve assembly opens at correct pressure
reading. Thermal relief valve is preset at factory to
open at 2050, ± 100 PSI. No further adjustment
should be necessary
Revision 3
5-59
MODEL 210 & T210 SERIES SERVICE MANUAL
| 5-172B.
DISASSEMBLY. (Refer to figure 5-26.)
NOTE
The relief valve assembly is preset by the
factory and normally will not require
disassembly. Refer to steps "h" and "i" of
paragraph 5-172D to determine if
disassembly or adjustment is necessary.
a. Remove nut (21) and adjustment screw (35 from
housing (20).
b. Remove spring (12), spring guide (19), balls (18),
and piston (13) from housing (20).
c. Loosen nut (21) and remove adapter (15) from
housing (20).
d. Remove poppet (17) and orifice (16) from adapter
(15).
5-172C. INSPECTION.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661 or equivalent) and dry with filtered
air.
b. Inspect all threaded surfaces for serviceable
condition and cleanliness.
c. Inspect all parts for scratches, scores, chips, cracks,
and indications of excessive wear.
ASSEMBLY ANDADJUSTMENT.
figure 5-26.)
| 5-172D.
(Referto
NOTE
When reassembling relief valve, install new
packing and internal parts lubricated with a
film of Petrolatum W-P-236, hydraulic fluid
MIL-H-5606, or Dow-Coring DC-7.
a. Installorifice(16)andpoppet 17)intoadapter(15).
(New packing must be installed on poppet.)
|
b. Install nut (21 and housing (20) on adapter (15).
c. Tighten adapter(15) into housing (20) and torque to
I 100-150 Ib-in (nut [211 must not contact housing [201
during torquing).
d. Tighten nut (21) against housing 20), and torque to
100-150 Ib-in.
e. Install one ball (18) into housing (20 so that it rests
on poppet (17). Install piston (13) into housing (20); then
install remaining ball (18) into end of piston t 13).
I f. Insertspring guide (19)andspring (12) into housing
(20) making sure that balls (18) and piston (13) remain in
correct position.
g. Turn adjustment screw (35) into housing (20) until
itjust contacts spring (12); then turn in one additional
| turn. Start nut (21) onto adjustment screw (35) and snug
against housing (20).
h. Connect a hydraulic pump with a flow rate of 0.5 to
0.7 gallons-per-minute, and a pressure gage with 2500
PSI capacity to relief valve. Apply pressure slowly to
insure that relief valve assembly opens and resets at the
following pressure readings:
1800 + 00 - 50 PSI
.........
OPEN
1300 PSI
RESET ...
(Leakage not to exceed 10 drops-per-minute.)
5-60
Revision 3
i. If adjustment of relief valve is necessary, turn
adjustment screw (35) in to increase pressure; back
adjustment screw out to decrease pressure. Tighten nut
(21) against housing (20) and torque to 100-150 Ib-in.
Recheck pressure adjustment.
5-173. DOOR SYSTEM THERMAL RELIEF VALVE.
(Refer to figure 5-26.) The relief valve is located in the
power pack assembly. The valve is preset at the factory
to open at 2050, ± 100 PSI. No further adjustment
should be necessary.
5-174. LANDING GEAR AND DOOR MANIFOLD
ASSEMBLIES. (Refer to figure 5-27.)
5-175. DESCRIPTION. The manifolds are mounted
on the power pack in the console. Refer to the schematic diagrams at the end of this Section for system
operation.
5-176. SOLENOIDS. The solenoids are mounted
on the top of the gear and door manifolds, and
should be disassembled, cleaned and reassembled
every 1000 hours or 5 years, and whenever the
solenoid is accessible.
5-177. DISASSEMBLY OF SOLENOID. (Refer to
figure 5-27.)
a. Cut safety wire and remove solenoid from
manifold.
b. Remove screws
c. Remove top,
d. Remove plunger.
e. Remove gland.
f.
Remove and discard packing,
5-178. INSPECTION AND CLEANING OF
SOLENOID COMPONENTS. Wash all parts in
solvent (Federal Specification P-S-661. or
equivalent) and dry with filtered air. If any parts
are found defective or worn. replace the entire
solenoid assembly. (Replace packing.)
5-179. ASSEMBLY OF SOLENOID. (Refer to
figure 5-27.)
a. Install new packing
b. Install plunger.
c. Install top
d. Install screws.
e. Install gland.
5-180. LANDING GEAR MANIFOLD. (Thru Serial
21062273.)
5-181.
DISASSEMBLY.
(Refer to figure 5-27.)
NOTE
As gear manifold assembly is removed
from body of power pack, transfer tube
(13) will fall free. Also, be careful of
spool (3), which is installed in top of
selector valve (4).
MODEL 210 &T210 SERIES SERVICE MANUAL
a. Remove packing (12) from bottom of manifold.
b. Remove packings (11) and (14) from transfer tube
(13).
c. Remove retainer (18) from gear manifold assembly.
Remove packings (19) from retainer.
NOTE
Retainer (18) is sealed in manifold
assembly with Loctite Hydraulic Sealant
or STA-LOK No. 550. or equivalent
sealant.
d. Remove AN316-4R nut (8) and screw (6).
e. Using a blunt tool or welding rod, push flow
valve spool (17) flow valve sleeve (24), spring (15)
and spring guide (26) through bottom of manifold
assembly.
NOTE
Use care to prevent damage to spring
guide (26), flow valve spool (17) or flow
valve sleeve (24).
f. Remove flowvalvespool (17) fromflowvalve sleeve
(24).
g. Remove packings (19) and (2) and back-up rings (20)
and (22) from flow valve sleeve (24).
h. Remove packing(16) from flow valve spool (17).
i. Remove spring guide (26) from spring (15), and
remove packing (25) and back-up ring (23) from spring
guide (26).
j. Cut safety wire and remove gear up-down solenoid
(1) from manifold. Remove packing (2) from gear updown solenoid (1).
k. Using a hook formed from brass welding rod, and
inserted into oil hole in selector valve (4), withdraw
selector valve from manifold.
CAUTION
Be sure that end of hook is not over 1/16inch long. Use care to prevent scratching
bore in manifold. Removal of selector
valve will be difficult due to friction
caused by packings.
1. Remove packings (5) from selector valve.
m. Remove spring (7).
5-181A. INSPECTION AND REPAIR.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661. or equivalent) and dry with
filtered air.
b. Inspect seating surfaces. They should have
very sharp edges. Seats may be lapped, if
necessary, with No. 1200 lapping compound.
c. Inspect all threaded surfaces for serviceable
condition and cleanliness. Clean sealant from
retainer threads.
Revision 3
5-60A/(5-60B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
d. Inspect all parts for scratches, scores, chips.
cracks and indications of excessive wear.
5-181B.
REASSEMBLY.
NOTE
When reassembling door manifold, install
new packings, back-up rings, and existing
threaded parts lubricated with a film of
Petrolatum W-P-236, hydraulic fluid
MIL-H-5606, or Dow-Corning DC-7.
a. Lubricate packings on selector valve (4).
b. Install packing in bottom of manifold.
c. Install spring (7) and selector valve (4) in
manifold,
NC7
Be sure spool (3) is installed in selector
valve (4) in position shown in Figure 5-27.
NOTE
assembly with Loctite Hydraulic Sealant
or STA-LOK No. 550, or equivalent
sealant.
d. Remove AN316-4R nut (8) and screw (6).
e. Using a blunt tool or welding rod, push flow
valve sleeve (4) and flow valve spool (11), spring (13)
and spring guide (16) through bottom of manifold body
(3).
NOTE
Use care to prevent damage to spring
guide (16), flow valve spool (11) or flow
valve sleeve (4).
f. Remove
flow valve spool (11) from sleeve (4).
g. Remove packings and back-up rings from
sleeve (4)
h. Remove packing from spool (11).
i. Remove packing and back-up ring from spring
guide (16)-
d. Install packing (2) on solenoid (1). Install
solenoid on manifold and safety wire as shown in
view AA
e Installscrew (6) and AN316-4R nut (8) in top
5 INSPECTION AND REPAIR.
o. manifold screw (6) and AN316-4R nut (8) in top
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661, or equivalent) and dry with
f. Install packing (25) and back-up ring (2) on
air.
filtered
spring guide (26).
b. Inspect seating surfaces. They should have
g. Install spring guide (26).
very sharp edges. Seats may be lapped, if
. Install spring (15).
necessary, with No. 1200 lapping compound.
i. Install packings (19 and 21) and back-up rings
(20 and 22) on flow valve sleeve (24).
c. Inspect all threaded surfaces for serviceable
j. Install spool (17) in sleeve (24); install
condition and cleanliness. Clean sealant from
assembly in bottom of manifold.
retainer threads.
k. Install packing (19) on retainer (18).
d. Inspect all parts for scratches, scores, chips,
1. Prime threads of retainer (18) with Grade T
cracks and indications of excessive wear.
Primer and seal with Loctite Hydraulic Sealant or
STA-LOK No. 550, or equivalent sealer.
5-183B. REASSEMBLY.
m. Install retainer (18).
a. Install screw (6) and AN316-4R nut (8) in top
n. Install packings on transfer tube (13).
of manifold.
o. Prior to installing manifold on body of power
b. Install packing (15) and back-up ring (14) on
pack. install transfer tube (13) in body of pack.
spring guide (16).
p. Refer to paragraph 5-184 for adjustment proc. Install spring guide (16).
cedures.
d. Install spring (13).
e. Install packings (1) and (2), and back-up rings
5-182. LANDING GEAR MANIFOLD. (Beginning
(5 and 7) on flow valve sleeve (4).
with Serial 21062274.)
f. Install packing (12) on spool (11).
g. Install spool (11) in sleeve (4): install assembly
5-183. DISASSEMBLY. (Refer to figure 5-28.)
in bottom of manifold.
h. Install packing (9) on retainer (10).
NOTE
i. Prime threads of retainer (10) with Grade T
Primer and seal with Loctite Hydraulic Sealant or
As gear manifold assembly is removed
STA-LOK No. 550, or equivalent sealer.
from body of power pack, transfer tube
j. Install retainer (10).
(18) will fall free.
k. Install packings (19) on transfer tube (18).
L Prior to installing manifold on body of power
a. Remove packing from bottom of manifold.
pack, install transfer tube (18) in body of pack.
b. Remove packings from transfer tube.
m. Refer to paragraph 5-184 for adjustment prcc. Remove retainer (10) from gear manifold
cedures.
assembly.
Revision 2
5-61
MODEL 210 & T210 SERIES SERVICE MANUAL
A
Safety wire solenoids (1)
and (27) after installing
on gear and door manifold assemblies.
Prime threads of retainer (18)
with Grade T Primer and seal
with Loctite Hydraulic Sealant
or STA-LOK No. 550,
equivalent sealant.2
SAFETY
WIRE
A
28
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
Gear Up-Down Solenoid
Packing
Spool
3
Selector Valve
Packing
Screw
4
Spring
AN 316-4R Nut
Gear Solenoid Assembly
Plug
5
Packing
Packing
Transfer Tube
6
Packing
7
Spring
7
8
Packing
Flow Vaive Spool
Retainer
Packing
Back-Up Ring
104
Packing
Back-Up Ring
Back-Up Ring
Flow Valve Sleeve
1
Packing
Spring Guide
Door Open-Close Solenoid
Retainer Ring
End Gland
31.
Back-Up Ring
32. Packing
33.
34.
35.
36.
37.
38.
39.
40.
41.
42.
43.
44.
45.
46.
47.
48.
46
31
30
3
31
32 ..
-
44
33
33
344
J
3'
3
·
26
23
42
3
38
4---1
24
50
12
13
Back-Up Ring
14
Piston
Door Manifold Assembly
Plug
15
Packing
Door Lock Valve
16
Packing
Packing
Packing
Transfer Tube
Plug
49.
Spring
50.
Packing
51.
52.
Selector Valve
53.
Spool
54.
Packing
A-A
View
29
22
/
O
40
/
1
17
Screw
Top
Plunger
Housing
Packing
Gland
'
8
GEAR MANIFOLD
ASSEMBLY
(Thru 21062273)
T
_.(
NOTE
52
DETAIL
During assembly, lubricate all packings
and back-up rings with a film of Petrolatum W-P-236, hydraulic fluid MIL-H5606, or Dow-Corning DC-7.
Figure 5-27. Gear Assembly Manifold and Door Manifold Assemblies
5-62
Revision 2
51
39
'
DOOR MANIFOLD
ASSEMBLY
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
5-184. ADJUSTMENT OF GEAR MANIFOLD ASSEMBLY
(Refer to figure 5-27 or 5-28.)
NOTE
With manifolds installed on power pack
and power pack installed on aircraft, if
main landing gear moves into the up or
down locks with sufficient force to jar the
aircraft, the flow control valve in the
landing gear manifold should be adjusted
in accordance with the following
procedures.
a. Jack aircraft in accordance with procedures
outlined in Section 2 of this manual, and attach
external power source.
b. Loosen AN 316-4R nut (8).
c. Back off screw (6) counterclockwise to
maximum snub position.
d. Rotate screw (6) clockwise to increase speed of
gear rotation and counterclockwise to slow speed
of gear rotation.
e. When desired setting has been achieved,
tighten AN 316-4R nut (8).
5-185. DOOR MANIFOLD ASSEMBLY. (Refer to
figure 5-27)
figure
5-186. DISASSEMBLY OF DOOR MANIFOLD.
(Refer to figure 5-27.)
NOTE
gland and piston.
5-187. CLEANING AND INSPECTION OF DOOR
MANIFOLD COMPONENTS.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661, or equivalent) and dry with
filtered air.
b. Inspect seating surfaces. They should have
very sharp edges. Seats may be lapped, if
necessary, to obtain sharp edges.
c. Inspect all threaded surfaces for serviceable
condition and cleanliness.
d. Inspect all parts for scratches, scores, chips,
cracks and indication of excessive wear.
5-188. REASSEMBLY OF DOOR MANIFOLD.
(Refer to figure 5-27)
NOTE
When reassembling door manifold, install
new packings, back-up rings, and existing
threaded parts lubricated with a film of
Petrolatum
-P-236, hydraulic flid
MIL-H-5606, or Dow-Corning DC-7.
a. Install new packings on end gland (29), piston
(34), selector valve (46) and transfer tube (42).
b. Install packings and door lock valve in bottom
of manifold.
c. Install spring (44) and selector valve (46) in
manifold.
As door manifold assembly is removed
from
from body
body of
of power
power pack,
pack, transfer
transfer tube
tube
(42) will
fall
free.
~valve
a. Remove packings (41) from transfer tube (42).
b. Remove packings from bottom of manifold,
(38).
and
and remove
remove door
door lock
lock valve
valve (38)
c. Remove spring (44).
d. Cut safety wire and remove solenoid (27);
remove packing (48) from solenoid.
e. Using a hook, formed from brass welding rod,
and inserted into oil hole in selector valve (46)
h.
withdraw selector valve from manifold
CAUTION
Be sure that end of hook is not over 1/16-189
inch long. Use with care to prevent
scratching
bore in
in manifold.
manifold. Removal of
of
scratching bore
selector
to
selector valve
valve will
will be
be difficult
difficult due
due to
f.
g.
Remove packings (45) from selector valve (46).
Remove spool (47) from selector valve.
h.
.
j.
k.
Remove
Remove
Remove
Remove
5-64
retainer ring (28).
end gland (29).
piston (34).
packings and back-up rings from end
Revision 2
NOTE
Be sure spool (47) is installed in selector
(46) in position shown in figure 5d.
d.
e.
28.)
Install packing (48) on solenoid (27).
Install packing (48) on solenoid (27).
Install solenoid on manifold and safety wire as
shown
in
view A-A.
f
Install piston (34) and end gland (29) in
f . Install piston (34) and end gland (29) in
manifold.
g. Install retainer ring (28).
Prior to installing manifold on body of power
pack, install transfer tube (42) in body of power
pack.
5-189. LANDING
GEAR
LANDING
GEAR HAND
HAND PUMP.
PUMP (Refer
(Refer to
to
figure 5-29.)
5-190. DESCRIPTION. The hand pump is located
in the cabin floor area between the pilot and copilot
seats. The pump supplies a flow of pressurized
hydraulic fluid to open the doors and extend the
landing gear if hydraulic pressure should fail.
5-191. REMOVAL OF LANDING GEAR HAND
PUMP. (Refer to figure 5-29.)
5-192.
DISASSEMBLY OF LANDING GEAR
HAND PUMP. (Refer to figure 5-29.)
MODEL 210 & T210 SERIES SERVICE MANUAL
5-195. LANDING GEAR POSITION SELECTOR
VALVE. (Refer to figure 5-30. ) A mechanical gear
position selector valve is located in the switch panel.
The pilot shuttles the valve mechanically when he
changes gear handle position. The handle must be
palled out prior to selecting gear position. Moving
the selector handle opens and closes ports in the
valve, enabling fluid under pressure to flow to the
various system components to retract or extend the
landing gear. A microswitch, mounted on the selector valve, is also actuated by movement of the selector handle and directs electrical current to the door
close solenoid and pump motor. Refer to the hydraulic system schematics at the end of this section for
switch circuitry.
5-195A. REMOVAL AND INSTALLATION. (Refer to
figure 5-30.)
a. Loosenjamnut(18) andremoveknob (19).
CAUTION
As hydraulic lines are disconnected, fluid
will leak. Precautions must be taken to
prevent excessive leakage, such as spreading drip cloths under fittings and capping
lines and fittings. Tag all electrical leads
to insure correct re-installation.
b. Disconnect four hydraulic lines routed to valve
and all electrical leads to micro-switch.
c. Remove screws attaching valve to instrument
panel.
SHOP NOTES:
5-66
Revision 3
d. Remove selector valve.
e. Reverse preceding steps to install gear selector
valve.
5-195B. DISASSEMBLY AND REASSEMBLY. (Refer
to figure 5-30.)
a. Remove cover (1), lock ring (3) and cap (4).
Thru 21063811, remove race (5) and bearing (6). Beginning with 21063812, remove washer (20).
b. Remove cotter pin (7), washer (8) and spring (9).
c. Pull rod (17) from disc (15); remove disc.
d. Remove pucks (11) and springs (12).
e. Reverse preceding steps for reassembly.
5-195C. INSPECTION OF PARTS. Replace packings
(10) and (16). Check valve for wear, foreign or abrasive materials. Disc (15) may be refaced (lapped) if
worn or abraded. Check rollers in bearings (6).
5-196. INSTALLATION OF LANDING GEAR
STRUT STEP. (Refer to figure 5-31.)
NOTE
Step is bonded to gear spring with Uralite
3121 or 3M EC-2216 adhesive.
a. Remove wheel, axle and fitting in accordance with
paragraph 5-52.
b. Mark position on inboard side of step that was
removed so that new step assembly will be installed
in as nearly the same position on the strut.
c. Remove all traces of the original bracket and
adhesive as well as any rust, paint or scale, with a
gear support, using shims (P/N 1241629) between outboard forging and landing gear support assembly.
The following shims are available from Cessna Parts
Distribution (CPD 2) through Cessna Service Stations.
manufacturer's instructions. Note pot life.
h. Spread a coat of mixed adhesive on bonding surfaces of strut and step assembly.
i. Slide new step up strut as far as it will go, then
use soft mallet to drive step to mark on strut. Be
sure step is level.
0.016 inch
1241629-1 ..............
0.025 inch
1241629-2 ..............
1241629-3.........
0.050 inch
0.071 inch
1241629-4 ..............
2. Use shims between downlock support assembly and outboard support assembly, to level wings and
assure that end points of main landing gear wheel axle
points are within ±0.25 inch.
CAUTION
It is important to install step in as nearly the
same location as old step. If step is not installed high enough on strut, during landing
gear retraction, step will contact top of strut
well wall.
j. Remove excess adhesive with lacquer thinner.
k. Allow adhesive to thoroughly cure according to
the manufacturer's recommendations before flexing
gear spring strut or apply loads to the step.
L Paint gear spring and step after curing is comm. Install wheel, axle and fitting.
5-197. RIGGING THROTTLE-OPERATED
MICROSWITCHES. (Refer to figure 5-32.)
Rigging procedures for sea level or turbocharged
aircraft are outlined in the figure.
5-198. RIGGING OF MAIN LANDING GEAR.
(Refer to figure 5 -34.)
5-68
Revision 3
NOTE
This
measurement may be made from a point
beneaththe wing main spar on the upper door
spring strut. Make measurements from corresponding points on the upper door sills.
Shim thickness between downlock support and
outboard support assembly shall not exceed
0.075 inch with a minimum thickness of 0. 025
inch for either main gear.
3. Before installing downlock hook (4),
adjustment screw (5), and arm assembly (7), adjust hook
MODEL 210 & T210 SERIES SERVICE MANUAL
SETTING THROTTLE SWITCHES
1. During night at 120 MPH (IAS), 2500', prop
control full forward for maximum RPM, and with
the gear and flaps up, mark the throttle control
position corresponding to the following manifold
pressures:
Model 210M
Model T210M_
12.0" * .5"
15.0" * 1.0"
REFER TO
SECTION
FOR CONTROL
LUBRICATION
2. Then adjust the gear warning horn throttle
switch on the ground to activate at the throttle
control position as marked in flight.
"For each 1000 feet above 2500' MSL, decrease
the manifold pressure at which the throttle control position is marked by 0.5 inches.
6
.
VIEW LOOKING AFT AND
OUTBOARD AT RIGHTHAND SIDE OF FIREWALL
1
1.
2.
3.
4.
5.
Switch Cover
Switch Cover
Spacer
Switch
Switch Spacer
Figure 5-32.
\
6. Switch Mounting Bracket
7. Arm Assembly
8. Gear Warning Cam
9. Fuel Pump Cam
10. Bushing
Rigging Throttle-Operated Microswitch
Revision
Revision 2
5-69
5-69
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
If it is planned to use the aircraft power system during rigging
procedures, outlined in the following paragraphs, the following
steps should be considered.
IMPORTANT POINTS CONCERNING
ELECTRO-HYDRAULIC SYSTEM INTERRELATIONSHIP
1. The electrical system is a 24-28 volt system (24 volt battery and 28 volt alternator).
The alternator is regulated to 27.7 volts, so bus voltage during engine operation will
be 27. 5 + 0.5 volts.
2.
The electro-hydraulic power pack motor requires a nominal 20 amps at 27.5 volts
during gear operation with starting current peaking out at 30 amps.
If the motor is operated in the shop on the ship's battery (engine not running), then
system voltage is only 22 to 24 volts during first and second gear cycles. It may be
even less if the ship's battery is old or partially discharged.
During landing gear system servicing, a power supply capable of maintaining 27.5
volts throughout the gear cycle must be used to augment the ship's battery.
3.
The power pack includes an electrically-driven pump and two electric solenoid
shuttle valves. These valves are normally energized during flight (gear retracted,
doors closed). The door valve is de-energized during the doors open and gear
cycling action. The door valve is re-energized at the end of the gearextensionor retraction cycle, causing the doors to close.
The pump motor is putting forth its maximum effort at about the same time the
door valve is energized. If the battery-alternator combination is not maintaining
27.5 volts, the gear valve may not shuttle. The doors remain open and the pump
continues to run.
The typical door solenoid will operate at 21.0 to 21.5 volts when hot. In a service
shop, when cycling the gear using a limited capability power source, the voltage
required to energize the door solenoid may not be developed.
5-70
Revision 2
MODEL 210 & T210 SERIES SERVICE MANUAL
setscrew (15) to stop hook 0.06, + 0.03, -0.20-inch
overcenter, as shown in figure 5-34.
4. Adjust downlock hook to clear inboard side
of gear pivot ear to a minimum of 0.06 inch.
NOTE
A spacer (P/N 1241614-1) is installed on each
side of the downlock arm assembly. Spacer
may be relocated to the inboard or outboard
side of the downlock arm assembly to obtain
the 0.06 inch clearance between hook assembly and the inboard of gear pivot ear. After
adjustment, both spacers MIGHT end up on
either the inboard or outside of downlock arm
assembly.
b. A new downlock actuator assembly is received
with a preassembled length of 12.45 inches, and the
three hydraulic ports in the same plane. Install actuator assembly, attaching it to fuselage structure
and downlock hook arm assembly.
c. With landing gear free, hydraulic pressure off, and
downlock system in position shown in figure 5-34, swing
gear into DOWN position and adjust adjustment screw
(5) as follows:
NOTE
To relieve hydraulic pressure, pull hydraulic
pump circuit breaker off, and move gear
selector switch up and down two or three
times.
adjustment screw. If hook (4) is under maximum
overcenter tolerance, green area of gage will contact
spacer on gear pivot, while red area will not make contact
with 0.050-inch diameter shoulder, as shown in figure
5-34. When hook (4) is on maximum overcenter
tolerance, both green and red areas will make contact. If
red area makes contact and green area does not, the hook I
setscrew (15) should be adjusted INWARD to bring
overcenter dimension to within tolerance.
3. Install 0.040-inch downlock gage (SE960) on
inboard side of hook (4) as shown in figure 5-34. If hook
(4) is over minimum overcenter tolerance, green area of
gage will contact shoulder, while red area will not make
contact with spacer.
4. When hook (4) is on minimum overcenter
|
tolerance, both green and red areas will make contact.
5. If overcenter tolerance is less than 0.040-inch,
the red area will make contact, while the green area will
not. If this condition exists, the next step is to determine
if the hook adjustment screw (5) is making contact with
the setscrew (15). This is accomplished by lifting the
landing gear spring upward off the hook (4) and checking
for possible rotation of the hook (4), by hand, with
hydraulic pressure off.
6. If a slight rotation is possible, hook setscrew
(15) is not contacting adjustment screw (5). If contact is
not being made, downlock actuator (25) will have to be
readjusted by backing off actuator's rod end one-half turn
at a time (one-and-one-half turn maximum adjustment)
until hook (4) is 0.040-inch or more overcenter and
contact is being made between setscrew (15) and
adjustment screw (4). If contact is being made, the hook
setscrew (15) should be adjusted outward to increase
overcenterness within tolerance.
1. If downlock locks, turn adjustment screw (5)
one-quarter turn OUT at a time until lock will not lock;
then turn in one-quarter turn and secure pin.
2. If downlock does not lock, turn adjustment
screw (5) one-quarter turn IN at a time until lock will
lock, and secure pin.
d. Readjust hook setscrew (15) to stop hook (4) 0.06,
+ 0.03, -0.02-inch overcenter as shown in figure 5-34.
e. When checking overcenter measurement of
downlock arm assembly, landing gear should be as
shown in figure 5-34, with nut, washer and spacer
removed, which retainl downlock arm assembly.
Use downlock overcenter gages (P/N SE960) to
determine if downlock hook assembly is still within
tolerance as shown on sheet 2 of figure 5-34. Use
gages as follows:
f. Now that hook adjustment screw (5) has been
adjusted, and hook setscrew (15) has been set to stop hook
at 0.06, + 0.03, -0.02-inch overcenter, check downlock
actuator rod end adjustment as follows:
1. Connect all hydraulic lines, fill system with
MIL-H-5606 hydraulic fluid and purge system of air
by cycling gear through several cycles.
NOTE
NOTE
NOTE
For correct rigging, hook setscrew (15) must
make contact with adjustment screw (5) and
green areas of both gages must contact as
shown in figure 5-34 for overcenterness to be
within tolerance.
Gages (P/N SE960) are available from Cessna
Parts Distribution (CPD 2) through Cessna
Service Stations.
Check fluid level in power-pack reservior
frequently during purging and rigging
procedures.
1. Remove nut, washer, and spacer which retain
arm assembly (7) to support assembly (3).
2. Install 0.090 downlock gage (SE960) on inboard
side of hook (4) as shown in figure 5-34. Upper portion of
gage should rest against head of pin attaching
2. Pull hydraulic pump circuit breaker off.
3. With gear in the down and locked position.
move the gear selector handle to the GEAR UP
position.
Revision 3
5-73
MODEL 210 & T210 SERIES SERVICE MANUAL
HEAD OF
MODEL 210 & T210 SERIES SERVICE MANUAL
position and note the actuation of main gear downlock hooks.
4. As soon as left downlock hook is actuated
to unlock left gear, move gear selector handle back
to "GEAR DOWN" position to simulate what would
occur if the pilot were to select gear down before the
gear was fully retracted.
5. If downlock hooks do not lock the gear in the
down position, check downlock system for misalignment.
g. With main gear in up-locked position, and system
pressure released, adjust uplock supports such that
ends of lock hooks are 0.92 inch inboard of lock hook
attach bolt. (Refer to figure 5-34. )
h. Adjust uplock system push-pull rods such that
when uplock latches are disengaged, both main gear
struts are released simultaneously and uplock studs
clear latches 0.15 inch minimum.
5-202. RIGGING OF NOSE GEAR LIMIT SWITCHES.
(Refer to figure 5-35.) The nose gear down indicator
switch is operated by an arm on the downlock mechanism. The nose gear up indicator switch is attached
to the uplock hook in the top of the nose wheel well.
After jacking the aircraft, adjust the switches as shown
in figure 5-35.
5-203. RIGGING OF NOSE GEAR SQUAT SWITCH.
The nose gear squat switch, electrically-connected to
the landing gear lockout solenoid, is operated by an
actuator, attached to the nose gear lower torque link.
Adjust the squat switch contacts to close when the
strut is between 0. 12 and 0. 25-inch from fully extended.
5-204. RIGGING RETRACTABLE STEP CABLE
ASSEMBLY. (Refer to figure 5-36.)
5-200. RIGGING OF NOSE LANDING GEAR.
(Refer to figure 5-35. )
Before working in landing gear wheel
wells, PULL-OFF hydraulic pump circuit breakers. Thru Serial 21062273,
the pump circuit breaker is located in
the circuit breaker panel, located immediately forward of the pilot's control
wheel. Beginning with Serial 21062274,
the pump circuit breaker is located In
the circuit breaker panel, located immediately forward of the left forward
doorpost. The hydro-electric power
pack system is designed to pressurize
the landing gear DOOR CLOSE system
to 1500 psi at any time the master switch
is turned on. Injury might occur to
someone working in wheel well area if
mater switch is turned on for any
reason.
NOTE
The nose gear downlock mechanism is
basically a claw hook at the end of the
piston rod end of the nose gear actuator.
a. Jack aircraft in accordance with procedures
outlined in Section 2 of this manual.
NOTE
The nose gear shock strut must be
correctly inflated prior to rigging the
nose gear. Refer to Section 1 of this
manual for correct nose shock strut
inflation.
b. The external claw locks on the nose gear
actuator shall completely engage lock pins without
drag, and crossbar shall rotate freely to indicate it
is not bearing on either side of slot in rod end.
Adjust rod end of actuator as required.
CAUTION
The piston rod is flattened near the
threads to provide a wrench pad. Do not
grip the piston rod with pliers. as tool
marks will cut the O-ring seal in the
5-201. RIGGING OF NOSE GEAR DOORS. Nose gear
door adjustments are accomplished with push-pull
rods as required to cause the doors to close snugly.
Doors must fair when the nose gear is fully retracted.
Link rods are to be adjusted so that the doors, when
in the open position, clear any part of the nose gear
assembly by a minimum of 0.25-inch during retraction. Trim outboard edge of nose gear doors so that
door-to-skin clearance is 0. 18-inch miniumum to
0. 21-inch maximum. Nose gear strut doors shall
fair when nose gear lock bushing is fully engaged with
uplock hook.
a. Rig nose gear in accordance with procedures
outlined in paragraph 5-200.
b. Rig nose gear doors in accordance with procedures outlined in paragraph 5-201.
c. Rig nose gear limit switches and nose gear squat
respectively.
d. While aircraft is still on jacks, extend landing
gear and disconnect strut door tie rods. DO NOT
DISTURB ROD ADJUSTMENT.
DISTURB
ADJUSTMENT.
e. Attach.ROD
retractable step assembly cable turnbuckle to spring clip at hook assembly on forward
end of nose gear actuator, if not previously attached.
f. Retract landing gear to up and locked positiong. Adjust retractable step assembly cable turnbuckle to hold cabin step in its best faired condition;
safety wire turnbuckle.
h. Extend landing gea and attach tie rods to strut
doors
NOTE
Install right-hand tie rod on outboard side
of eyebolt only, when connecting nose gear
strut doors. Left-hand tie rod should be
installed in normal manner.
i.
5-78
Revision 2
Remove aircraft from jacks.
MODEL 210 & T210 SERIES SERVICE MANUAL
SQUAT
Thru Serial 21062273
SWITCH
F-CDId
UPLOCK
~/,^
SWITCHES
B
F.GOn
FGd
,
/
G
G--
'A'GE4
I
l
-'h
^
OOF.GO22K-coo
F--D^rd-OO.GD2I1-d
US
F-GD0 NOSE
IS
G.Dc2
LEFT
5HT
I F.CDI
9
1 L^
tGEA
-----
F.CE5
DOWNLOCK SWITCHES
MOTIOR
I
I.II
,
I
F-GDIJ
F.Col(GDIS
I
IDOWN
F-GDI
,
GD7FCF.GD5
30F.G2
6
I1 ,__ -'--
-rf''.Goo
/l'llll
te
--
\
i
*
5-86
Revision 2
MAJLUlN-
DOORS
mittEAI
f .Hi
-F-GDI0.
F.GD-(
.,.. , I
A*. UP
FILE11 1S11
GEAR
|SUCllON
UP *
Nl-mNi
tt UC
FF-G°3 S
tl
--
STA
GEAR DOWN C
-
E
,234
1.D2 6 S I
IN-NEPLUG
LI 34 S (JILOCATED IN CONSOLE)
WHITE
SILEEINI-D
*
-
.I
CUlIP
1----
, t.l,
F.C 12
'-O--.'
---
rU
OR
I I lt
RESE
.M
=
DOORS OPEN
*
11
INN
BBIU
TIC
i | PRE1Si
SSUR E
-iMllKll
CLOSED
EAR IYIIAIItl PiOeR PACK
nu) DOORS OPEN
+ GEAR DOWN IAMB
I AGEAR UP
* DOORS CLOSED
STAIIC
11111
PSTA HL
PRESSURE
SUCTION
GEAR DOWN - DOORS CLOSED & MOTOR TURNING OFF
Figure 5-37.
5-86
Revision 2
Hydraulic and Electric System Schematic
(Sheet 6 of 7)
I
TIATII
PRESSURE
TFLOW
STATIC
MODEL 210 & T210 SERIES SERVICE MANUAL
Thru Serial 21062273
SQUAT
SWITCH
,.OrOO,
IND
OSE
RIGHT
S-G'7-
0&)
-
-oeC
Q.D20
C
'
,.,I
f.-D2-3
LEFT
--
-03
s
O
°OW"NLOCK SWITCHESI_
MOTOR IWIICNES
w
t
la-GD
DOWN
*GCD13
I
Lf.-GD7
_
.--
I
~~3O~~l
KUanI
^
--
0
--
*2
HITE
11"1r'^Ee'~ *Ul~~
tu l
-;-o
J
i
,io
flaa
e5i3
l
lr
7.
H
a
c a
El
!
4f 526 1 IN-LINE PLUG
3 4 5jN(LOCAIED IN CONSOLE)
cVrrrrr
1l
IUP* DOOMIR 11 CL
a
( h em R llK
GOr~-00QF-!d
f .)RT
GEAR
-2
cIct
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i
U~V
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;
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Ta MAS-"ISWHIT
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I I _ E-CfleS
UU5.Ul3
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Fig
U,
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F
IMM
UllE
SlE
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Ul
IlllBS111l
pm[IE
suzu
_
5_87
llPlAK
itl
uuGL
unnn
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i
_11
SYSTEM
u
~~GEiAR DOWN
*
Ku
MICI
COMPLETE (AIRCRAFT
Figure 5-37.
GEAR UP
MASTER
In
1
DOORS OPEN
* DOORS CLOSED
SWITCH
OFF)
Hydraulic and Electric System Schematic
CODE l
PRESURE
iSTATIC
PRESSUREz|
I
ON
a
RETURN
TAC
OLIGHTON
(Sheet 7 of 7)
Revision 2
5-87
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 5A
LANDING GEAR. BRAKES AND HYDRAULIC SYSTEM (BEGINNING WITH 1979 MODELS)
WARNING
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the propeller,
since a loose or broken wire or a component
malfunction could cause the propeller to rotate.
NOTE
This section covers 1979 and ON models, and was
added to avoid the confusion of serialization caused
by major changes in the aircraft hydraulic system.
However, Section 5 contains information which is
also applicable tothese models. To avoid repetition, the reader is referred back to Section 5 for
this information.
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
LANDING GEARSYSTEM .........
Description ................
Trouble Shooting ............
Hydraulic System Leak Check
Power Pack
................
1I17/5A-3
1I7/5A-3
1I18/5A-4
1I22/5A-8
Reassembly ............
Adjustment ............
Emergency Hand Pump .......
Description ...............
1J15/5A-18
1J16/5A-18A
1J16/5A-18A
1J16/5A-18A
1I23/5A-9
Removal and Installation
1J17/5A-18B
Description ...............
1I23/5A-9
On-Aircraft Hydraulic Power
Pack Operational Checks . 1I23/5A-9
Removal .............
1I23/5A-9
Disassembly ..............
1I24/5A-10
Inspection .............
1I24/5A-10
Reassembly ...........
1I24/5A-10
Installation ...............
1J1l/5A-14
Primary and Thermal Relief
Valve Assemblies ........
1J12/5A-15
Bench Check of Primary and
Thermal Relief Valves . 1J12/5A-15
Removal ...............
1J12/5A-15
Disassembly ...........
1J12/5A-15
Inspection .............
1J12/5A-15
Assembly and Adjustment 1J12/5A-15
Installation ............ 1J13/5A-16
Pressure Switch ...........
1J13/5A-16
Description ............
1J13/5A-16
Removal (Thru 21063964
plus 21063973) ........ 1J13/5A-16
Disassembly ...........
1J13/5A-16
Inspection and Repair ...
1J14/5A-17
..
Disassembly ..............
1J17/5A-18B
Inspection and Repair ......
1J17/5A-18B
Reassembly ...............
1J17/5A-18B
Landing GearSelector Valve ..
1J17/5A-18B
Description ...............
1J17/5A-18B
Removal and Installation
1J18/5A-19
Disassembly and Reassembly 1J18/5A-19
Inspection and Repair ...... 1J19/5A-20
Rigging Throttle-Operated
Warning Horn Microswitch
1J19/5A-20
Main Landing Gear .......... 1J19/5A-20
Description ............... 1J19/5A-20
TroubleShooting ..........
J21/5A-22
Removal ..................
1J21/5A-22
Installation ............... 1J21/5A-22
Rigging ..................
1K1/5A-26
Rigging Main Gear Down
Limit Switches ........... 1K2/5A-27
Rigging Main Gear Up
LimitSwitches ...........
1K2/5A-27
Main Wheel and Tire ......
1K5/5A-30
Description ............
1K5/5A-30
Balancing and Alignment
Main Wheel and Axle ...
1K5/5A-30
1K5/5A-30
1J15/5A-18
1J15/5A-18
Main Gear Actuator .......
Removal ...............
Disassembly ...........
Inspection ..............
1K5/5A-30
1K5/5A-30
1K5/5A-30
1K6/5A-31
1J15/5A-18
1J15/5A-18
1J15/5A-18
Parts Repair/Replacement
Reassembly ............
Installation ............
1K6/5A-31
1K6/5A-31
1K6/5A-31
Reassembly ............
1J14/5A-17
Adjustment ............
Installation ............
Removal (21063965 thru
21063972 and 21063974
& on) ................
Disassembly ..........
Inspection and Repair ...
Revision 3
5A-1
MODEL 210 & T210 SERIES SERVICE MANUAL
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
Main Gear Pivot Assembly .
Removal ...............
Inspection and Repair ...
Installation ............
Gear Position Indicator ....
Switches ..............
Description ............
Main Gear Downlock Actuator
Description ............
Main Gear Strut Step ......
Description ............
Removal ...............
Installation ............
Nose Gear System ............
Description ...............
Operation ................
Trouble Shooting ..........
Removal of Nose Gear
Assembly ..............
Shimmy Dampener ........
Torque Links .............
Squat Switch .............
Nose Gear Downlock
Mechanism ..............
Nose Gear Actuator .......
5A-2
Revision 3
1K6/5A-31
1K6/5A-31
1K6/5A-31
1K6/5A-31
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K7/5A-32
1K8/5A-33
1K8/5A-33
Nose Gear Door System ....
Description ............
Removal and Installation
Nose Wheel Steering System
Description ............
Rigging Nose Landing Gear
Rigging Nose Gear Down
Limit Switch ............
Rigging Nose Gear Up
Limit Switch ............
Rigging of Nose Gear Squat
Switch ..................
Rigging of Nose Gear Doors .
Final Landing Gear Systems
Check ...................
Nose Wheel and Tire .......
Brake System ................
Brake Master Cylinder .....
Description ............
Removal ...............
Disassembly ...........
Inspection and Repair ...
Reassembly ............
Installation ............
Parking Brake System .....
1K8/5A-33
1K8/5A-33
1K8/5A-33
1K8/5A-33
1K8/5A-33
1K8/5A-33
1K8/5A-33
1K8/5A-33
1K8/5A.33
1K8/5A-33
1K8/5A-33
1K12/5A-37
1K12/5A-37
1K12/5A-37
1K12/5A-37
1K12/5A-37
1K12/5A.37
1K12/5A-37
1K12/5A-37
1K12/5A.37
1K12/5A-37
MODEL 210 & T210 SERIES SERVICE MANUAL
WARNING
When performing any inspection or maintenance that requires turning on the master
switch, installing a battery, or pulling the
propeller through by hand, treat the propeller as if the ignition switch were ON. Do
not stand, nor allow anyone else to stand,
within the arc of the propeller, since a loose
or broken wire, or a component malfunction,
could cause the propeller to rotate.
5A-1.
LANDING GEAR SYSTEM.
5A-2. DESCRIPTION. Retraction and extension of
the landing gear is accomplished by a hydraulicallypowered system, integrated with electrical circuits
which help control and indicate gear position. Retraction and extension of the landing gear incorporates a nose gear actuator and two main gear actuators. The main gear actuators control the main
gear struts through a sector gear arrangement. The
nose gear doors are mechanically-operated. The
doors are closed with the gear retracted and are open
with the landing gear extended. The main gears have
no doors. Hydraulic fluid is supplied to the landing
gear actuating cylinders by an electrically-powered
power pack assembly, located inside the center console. The hydraulic reservoir is an integral part of
the power pack assembly. Gear selection is accomplished manually by moving a gear selector handle,
located immediately left of center, in the switch
panel. It is necessary to pull out on the gear selector
to move the handle up or down. For emergency ex-
tension of the gear, the selector handle must be in the
DOWN position before the hand pump will energize
the system. A pressure switch is mounted on the
pump body. This switch opens the electrical circuit
to the pump solenoid when pressure in the system increases to approximately 1500 psi. The pressure
switch will continue to hold the electrical circuit open
until pressure in the system drops to approximately
1000 psi. This will occur whether the gear selector
handle is in either the UP or DOWN position. During
a normal cycle, landing gear extended and locked can
be detected by illumination of the gear DOWN indicator (green) light. Indication of gear retracted is
provided by illumination of the UP indicator (amber)
light. The nose gear squat switch, activated by the
nose gear, electrically averts inadvertent retraction
whenever the nose gear strut is compressed by the
weight of the aircraft. Beginning with 1983 models,
the up indicator (amber) light is replaced with a GEAR
UNSAFE indicator (red) light. The GEAR UNSAFE
(red) light is on anytime the gear is in transit (retract or extend), or whenever system pressure drops
below 1000 PSI with the safety (squat) switch closed.
NOTE
It is possible to have the red and green lights
on momentarily at the same time after the
completion of the extend cycle, or when rotating during takeoff. However, if both stay
on after the completion of the extend cycle,
or if the red light stays on longer than 5 to
7 seconds during the retract cycle, a malfunction has occurred.
SHOP NOTES:
5A -3
MODEL 210 & T210 SERIES SERVICE MANUAL
5A-3.
TROUBLE SHOOTING.
Just because this chart lists a probable cause, proper checkout procedures cannot be deleted and the replacement of a part is not necessarily the proper solution to the problem. The mechanic should always look for obvious problems such as loose or broken parts, external leaks, broken wiring, etc. To find the exact cause of a
problem, a mechanic should use a hand pump, pressure gage and a voltmeter to isolate each item in the system.
nydraulic fluid william if' air is pumped into system, causing fluid to be blown overboard thru pack vent line.
The problems listed are all with the systems controls in their normal operating position: Master switch ON,
hydraulic pump breaker IN and landing gear breaker IN. During landing gear system servicing, a power
supply capable of maintaining 27. 5 volts throughout the gear cycle must be used to augment the ship's battery.
CAUTION
Prior to using Hydro-Test unit with power pack, remove and dry off
filler plug and dipstick. Adjust cap tension so that no movement
of cap is apparent. Failure to accomplish these procedures could
result in filler cap coming loose from power pack.
TROUBLE
MOTOR PUMP WILL NOT
OPERATE GEAR BUT
EMERGENCY HAND PUMP
WILL OPERATE GEAR.
REMEDY
PROBABLE CAUSE
Low voltage (in flight).
Check alternator and wiring.
Fluid level low in reservoir.
Refill reservoir.
Motor pump failure.
Replace pump.
Faulty check valve
Replace valve
NOTE
Motor and pump are not repairable and must be replaced.
PUMP OR EMERGENCY PUMP
WILL NOT BUILD PRESSURE
IN SYSTEM.
5A -4
Pump frozen.
Remove motor and coupling
from top of power pack and
replace pump.
Broken pump or motor drive
shaft or coupling,
Remove motor and pump from
top of power pack and replace
motor, pump and coupling.
If motor was not turning,
check wiring and motor.
Check motor for loose or broken
connections; check for frozen
pump or coupling. Check
circuit breaker in pedestal.
Bad pump shaft seal.
Replace pump.
External leakage around top
of pump assembly
Remove motor and pump assemblies from top of power pack and
replace upper packing and/or
back-up rings
Air lock in pump (new pack
installation or pump replacement.
Remove filter and intermittenly
bump start switch until fluid flows.
Replace filter.
No fluid in reservoir.
Refill reservoir.
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont)
TROLBLE
PUMP OR EMERGENCY PUMP
WILL NOT BUILD PRESSURE
IN SYSTEM. (Cont).
HAND PUMP DOES NOT BUILD
PRESSURE, BUT ELECTRIC
PUMP OPERATES PROPERLY.
LANDING GEAR OPERATION
EXTREMELY SLOW.
POWER PACK EXTERNAL
PROBABLE CAUSE
Broken hydraulic Line.
i Check for evidence of leakage
and repair or replace line.
Flush out system and refill
reservoir.
Bad O-ring actuator
piston; O-ring left out
after repair.
Disconnect line upstream from .
actuator and check for pressure.
Perform this check for all
actuators in system.
Bad O-ring on gear
control valve.
Replace O-ring.
Thermal relief valve stuck open.
Replace valve.
Check valve in hand pump
sticking.
Inspect check valve.
Defective hand pump outlet check
valve.
Replace valve.
Main gear or downlock actuator
O- ring leaking.
Disassemble actuator and
replace O-rings.
Filter in outlet check valve improperly positioned in filter
body, or seal between filter
and check valve improperly
positioned.
Replace seal and position
filter in retainer with
Petrolatum.
Downlock rod adjustment
incorrect (mainly LH rod).
Adjust rod end to lengthen
actuator one turn.
Pump failure.
Replace pump.
Low voltage in electrical system.
Check alternator and wiring.
Replace pump motor.
Pump motor brushes worn.
Fluid leak in gear line.
Locate and repair or replace
broken line or fitting.
Excessive internal power pack
leakage.
Remove and repair or replace
power pack.
Static seals (all fittings).
LEAKAGE.
GEAR DOWN-LOCK WILL NOT
RETURN TO FULL-LOCK
REMEDY
Remove and replace O-rings
and/or back-up rings as
required. Check tubing
flares for leaks.
Reservoir cover.
Remove power pack and remove
cover; replace seals.
Binding in spring and
tube assemblies.
Check operation to locate
binding and eliminate.
POSITION.
5A -5
MODEL 210 & T210 SERIES SERVICE MANUAL
5A-3.
TROUBLE SHOOTING.
TROUBLE
LANDING GEAR FAILS TO
RETRACT.
GEAR RETRACTION OR EXTENSION EXTREMELY SLOW.
PUMP MOTOR STOPS BEFORE
GEAR IS RETRACTED.
PUMP MOTOR STOPS BEFORE
GEAR IS EXTENDED.
5A -6
PROBABLE CAUSE
REMEDY
Hydraulic pump motor circuit
breaker open.
Reset. determine cause for opening. Repair or replace components as necessary.
Instrument panel gear indicator
circuit breaker open.
Reset breaker. Determine cause
for tripped breaker.
Hydraulic pump motor circuit
wires disconnected or broken.
Repair or replace wiring.
Instrument panel gear indicator
circuit wires disconnected or
open.
Repair or replace wiring.
Nose gear squat switch inoperative.
Install new switch.
Pressure switch defective.
Install new switch.
Hydraulic pump motor solenoid
defective.
Install new solenoid.
Hydraulic pump motor ground.
Check for ground.
Hydraulic pump motor defective.
Replace motor.
Reservoir fluid level below
operating level.
Fill reservoir with hydraulic
fluid.
Battery low or dead.
Check battery condition. Install
new battery.
Reservoir fluid level below
operating level.
Fill reservoir with hydraulic
fluid (Refer to Section 2).
Restriction in hydraulic system.
Isolate and remove restrictions.
Hydraulic pump motor circuit
breaker open.
Reset, determine cause for
opening. Repair or replace
components as necessary.
Instrument panel gear indicator
circuit breaker open.
Reset circuit breaker. Determine
cause of tripped circuit breaker.
Pressure switch out of adjustment.
Remove, adjust or install new
switch.
Restriction in hydraulic system,
allowing pressure to build up
and shut off pump motor before
gear is retracted.
Isolate and determine cause.
Remove restriction.
Hydraulic pump motor circuit
breaker open.
Reset, determine cause for opening. Repair or replace components as necessary.
Instrument panel gear indicator
circuit breaker open.
Reset circuit breaker. Determine
cause of tripped circuit breaker.
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont.)
TROUBLE
PUMP MOTOR CONTINUES
TO RUN AFTER GEAR IS
FULLY RETRACTED OR
EXTENDED.
PUMP MOTOR CYCLES
EXCESSIVELY AFTER
GEAR IS RETRACTED.
GEAR DOES NOT FULLY
RETRACT, BUT PUMP
MOTOR CONTINUES TO
RUN.
LANDING GEAR FAILS
TO EXTEND.
PROBABLE CAUSE
REMEDY
Pressure switch defective.
Install new switch.
Pressure switch out of adjust.
Remove, adjust or install
new switch.
Hydraulic pump motor
solenoid defective.
Install new solenoid.
Internal leakage in system.
Check actuators for internal
leakage. Repair or install
new actuators.
External system leakage.
Check all lines and hose for
leakage. Repair or install
new parts.
Power pack relief valve out of
adjustment.
Disassemble and repair or
replace valve assembly.
Hydraulic motor solenoid
defective.
Install new solenoid.
Pressure switch out of adjustment.
Remove, adjust or install
new switch.
Internal leakage in system..
Check actuators for internal
leakage. Repair or install
new actuators.
External system leakage.
Check all lines and hose for
leakage. Repair or install
new parts.
Internal leakage in system.
Check actuators for internal
leakage. Repair'or install
new actuators.
Reservoir fluid level below
operating level.
Fill reservoir with hydraulic
fluid (Refer to Section 2).
Battery low or dead.
Check battery condition.
Install new battery.
Hydraulic pump motor circuit
breaker open.
Reset, determine cause for
opening. Repair or replace
components as necessary.
Instrument panel gear indicator
circuit breaker open.
Reset circuit breaker. Determine cause of tripped
circuit breaker.
Hydraulic pump motor circuit
wires disconnected or broken.
Repair or replace wiring.
Hydraulic pump motor solenoid
defective.
Install new solenoid.
5A -7
MODEL 210 & T210 SERIES SERVICE MANUAL
TROUBLE SHOOTING (Cont. )
TROUBLE
LANDING GEAR FAILS
TO EXTEND (cont).
PROBABLE CAUSE
REMEDY
Hydraulic pump motor ground.
Check ground.
Hydraulic pump motor defective.
Replace motor.
Reservoir fluid level below
operating level.
Fill reservoir with hydraulic
fluid (Refer to Section 2. )
Nose gear contacts stop bolts.
Adjust stop bolts to obtain
proper clearance. (Refer
to paragraph 5A-87).
RH GEAR UNLOCKS BUT
LH GEAR WILL NOT
UNLOCK.
Improper setting of RH downlock
actuator rod.
Check rigging procedures
outlined in this Section.
BOTH RH AND LE MAIN GEAR
UNLOCK BUT ONLY NOSE
GEAR WILL RETRACT.
Improper setting of LH downlock
actuator rod.
Check rigging procedures
outlined in this Section.
MOTOR PUMP WILL NOT
TURN ON BY WORKING
SELECTOR SWITCH. HAND
PUMP WILL PUT GEAR DOWN.
Defective pressure switch
circuit.
Check circuit continuity.
Check switch adjustment
SET SCREW ON CAM NOT EXCheck washers under bolt
TENDED ENOUGH FOR GEAR TO on downlock arm assembly.
MOVE CAM OVER CENTER.
Add AN960-10 washer under
bolt downlock arm assembly
MAIN GEAR WILL NOT LOCK
OVER CENTER.
Main gear not centered in
support.
Rerig saddle per rigging
instructions.
MALFUNCTION OF GEAR
INDICATOR LIGHTS.
1.
2.
Check ground wire for proper
connection.
15A-3A.
Both lights on at same time.
Light will change from green
to amber or in reverse when
gear control switch is moved.
HYDRAULIC SYSTEM LEAK CHECK. (Refer
to figure 5A-2.)
a. Jack aircraft in accordance with procedures in
Section 2 of this manual,
b. To relieve system pressure, pull the GEAR PUMP
circuit breaker to OFF, move the gear selector handle to
UP, and move back to the DOWN position.
c. Install a 0-2000 PSI gage at the tee (Index 28, figure
5A-3) on the left side of the power pack.
d. Push the GEAR PUMP circuit breaker to the ON
position, turn ON the master switch, and move gear
selector handle to the UP position.
e. Monitor pressure gage, after retraction cycle is
complete, for pressure bleed down.
f. If bleed down occurs, it can be an internal or
external leak anywhere in the system.
5A-8
Revision 3
NOTE
When any line is disconnected, be prepared
for fluid leakage.
g. Disconnect the return line from the gear selector. If
fluid comes from the selector, the internal leak is in the
system.
h. If no leak-by is found, it can be assumed there is an
internal leak in the power pack. If leak is found, proceed
to step "j." Reconnect the return line.
i. Power pack internal leakage can only be attributed
to a bad thermal relief valve, check valve, or check valve
O-ring. The only way to isolate part that is leaking is to
systematically replace the check valve O-ring, check
valve, and then thermal relief valve. Repeat leak test
after replacement of each part to ensure leak correction.
MODEL 210 &T210 SERIES SERVICE MANUAL
j. Remove gear DOWN line from the selector. If
fluid comes from the line, one or more of the gear
actuators is leaking. To locate the leaking actuator,
disconnect the return line from each actuator. the
leaking actuator will have fluid draining from the
actuator port. Following the appropriate paragraphs
in this section remove, overhaul and reinstall the
actuator.
k. Reconnect gear down line to the selector,
1. Recheck all lines that were disconnected for
security.
m. Lower the landing gear. Following the procedures in step "b" relieve the system pressure.
n. Remove the pressure gage from the service tee.
o. In accordance with the procedures in Section 2
of this manual replenish the power pack reservoir
with MIL-H-5606 hydraulic fluid and bleed the system.
p. Remove aircraft from jacks.
5A-4.
POWER PACK.
(Refer to figure 5A-3.)
5A-5.
DESCRIPTION. The hydraulic power pack,
located in the pedestal, is a multi-purpose control unit. It
contains a hydraulic reservoir, valves, an electricallydriven motor, and the pump. An emergency hand pump,
located between the pilot's and copilot's seats, uses
reservoir fluid to permit manual extension of the landing
gear.
NOTE
The hydraulic power pack primary relief
valve, thermal relief valve, and pressure
switch can be operationally checked on the
aircraft without power pack removal from the
aircraft or disassembly. Refer to paragraph
5A-5A for specific instructions. Refer to
paragraph 5A-11A for primary and thermal
relief valve bench check instructions if the
power pack is removed from aircraft.
5A-5A. ON-AIRCRAFT HYDRAULIC POWER PACK
OPERATIONAL CHECKS. (Refer to figure 5A-3.)
The primary and thermal relief valves and pressure
switch should be pressure checked each 100 hours. They
can be operationally checked without removal from
aircraft. For bench check instructions after removal from
power pack, refer to paragraph 5A-11A.
NOTE
Checks are to be performed with external
power set at 28.5 volts.
a.
Primary Relief Valve.
(1) Jack aircraft in accordance with procedures
outlined in Section 2.
(2) Remove cap and install pressure gage at tee (28)
fitting on left side of power pack.
(3) Pull landing gear circuit breaker.
(4) Select landing gear handle to DOWN position.
(5) Install 18 gage (minimum)jumper wire between
buss side of contactor and small terminal on pump
motor contactor (to energize coil).
(6) Push landing gear circuit breaker in; power pack
should run; monitor pressure.
(7) Primary Relief valve should open at 1800 PSI,
+0or-50PSL
(8) After check is complete, remove pressure from
system, remove pressure gage, install cap on tee (28),
pull landing gear circuit breaker, remove jumper wire,
push landing gear circuit breaker back in, and return
system to original configuration.
b. Thermal Relief Valve.
(1) With aircraft on jacks and pressure gage
installed at tee (28) fitting on left side of power pack,
pull landing gear circuit breaker.
(2) Select landing gear to DOWN position.
(3) Etend emergency gear pump handle.
(4) Pump emergency gear pump handle and monitor
pressure. Thermal relief valve should open at 2200
PSI, -0 or + 50 PSI.
(5) After check is complete, remove pressure from
system, remove pressure gage, and install cap on tee
(28).
(6) Push in landing gear circuit breaker, and return
system to original configuration.
c. Pressure Switch.
(1) With aircraft on jacks and pressure gage
installed at tee (28) fitting on left side of power pack,
pull landing gear circuit breaker.
(2) Select landing gear UP and DOWN several times
to relieve pressure in landing gear system.
(3) Select landing gear UP, and push in landing gear
circuit breaker.
(4) After gear raising cycle is complete, check
pressure. Pressure should be 1500 PSI.
(5) Select gear DOWN. After gear lowering cycle is
complete, pressure should be 1500 PSL
(6) After check is complete, remove pressure from
system, remove pressure gage, install cap on tee, and
return system to original configuration.
5A-6.
REMOVAL.
(Refer to figure 5A-3.)
a. Jack aircraft in accordance with procedures
outlined in Section 2 of this manual.
b. Turn master switch OFF and place gear selector
handle in a neutral position to relieve system pressure. After 15 seconds, return gear selector handle
to DOWN position.
NOTE
As hydraulic lines are disconnected or removed, plug or cap all openings to prevent
entry of foreign material into the lines or
fittings.
c. Remove front seats and spread drip cloth over
front carpet.
d. Remove decorative cover from pedestal as outlined in Section 9 of this manual.
e. Remove upper panel assembly from aft face of
pedestal.
f. Remove screws attaching indicator assembly at
top of pedestal; remove indicator assembly.
g. Remove four bolts attaching wheel and gear box
assembly; remove wheel and gear box assembly.
Revision 3
5A-9
MODEL 210 &T210 SERIES SERVICE MANUAL
h. Loosen idler sprocket assembly by loosening
bolt and sliding sprocket inboard in slot.
i. Disconnect chain at connecting link.
j. Remove left-hand and right-hand chain guards.
k. Allow chain to remain on gimbal assembly in
lower pedestal area.
1. Position gallon container under drain elbow at
right-hand side of pedestal.
m. Remove cap from elbow and attach drain hose.
n. Using hand pump, drain reservoir fluid into
container.
o. Disconnect and cap or plug all hydraulic lines
at power pack.
p. Disconnect wiring at pressure switch.
q. Remove three mounting bolts, one at the forward side of power pack, and two, attaching power
pack bracket to sides of pedestal.
r. Remove power pack and bracket from pedestal
as a unit.
NOTE
It should not be necessary to disturb studs
on left and right sides of pedestal to remove power pack.
5A-7. DISASSEMBLY. (Refer to figure 5A-3.)
a. Remove bolts (24), washers (25), and packing
(26) from reservoir (1).
b. Remove reservoir (1) from body assembly (19).
NOTE
If reservoir (1) will not disengage from body
assembly (19), install a capped fitting in the
pressure and return openings of the power
pack assembly and attach an air hose to vent
fitting at top of body assembly (19). Apply air
pressure (not to exceed 15 PSI, reservoir proof
pressure), and remove reservoir (1). A strap
clamp is not recommended as clamp may
damage reservoir (1).
c.
Remove packing (20) from body assembly (19).
NOTE
normally
not required. Refer to applicable
paragraphs for specific instructions regarding relief valves. Before removal, tag each
relief valve (primary) or (thermal) to ensure
correct reinstallation.
5A-10
Revision 3
d. Cut safety wire and remove relief valve assemblies (5) and (23) from body assembly (19).
e. Remove dipstick (15) and fluid filter screen (16)
from body assembly (19).
f. Remove retainer (12), self relieving check valve
filter assembly (11), back-up ring (13), packing (14),
packing (10) and check valve (9) from body assembly
(19).
NOTE
If check valve (9) will not fall from hole in
body assembly (19), place a drift or punch
made of soft material into the pressure
opening of body assembly (19) and tap spacer
from body assembly (19).
g. Remove pressure switch (17) and packing (18) from
body assembly (19).
I
h. Cutsafety wire, serial 21064588 and on, and
remove bolts (4) attaching hydraulic pump (6) to body
assembly (19), and remove pump (6) and coupling (8)
from body assembly (19). Remove packings (20) and (22).
i. Remove motor assembly from body assembly (19) by
removing attaching bolts (4).
5A-8. INSPECTION. (Refer to figure 5A-3.)
a. Wash all parts in cleaning solvent (Federal
Specification P-S-611, or equivalent) and dry with
filtered air.
b. Inspect all threaded surfaces for serviceable
condition and cleanliness.
c. Inspect all parts for scratches, scores, chips, cracks,
and indications of excessive wear.
d. Clean to ensure that all screens and filters are
completely clean and undamaged.
5A-9.
REASSEMBLY. (Refer to figure 5A-3.)
NOTE
During assembly, lubricate new packings,
back-up rings, and threaded surfaces with a
film ofPetrolatum W-P-236, hydraulic fluid
MIL-H-5606, or Dow-Corning DC-7.
a. Using new packing (22), install hydraulic pump (6)
and coupling (8) into body assembly (19) with bolts (4).
MODEL 210 AND T210 SERIES SERVICE MAN UAL
MODEL 210 &T210 SERIES SERVICE MANUAL
SELECTOR·
*POWER PACES.--
.
.-
..
Figure 5A-2.
Tanding Gear System Component Locator
Revision 3
5A-11D/(5A-12 blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
*
THRU SERIAL 21063964 PLUS
SERIAL 21063973
*
SERIAL 21063965 THRU SERIAL 21063972
AND 21063974 & ON
14
15
* 21062955 THRU 21064587
* REFER TO SERVICE INFORMATION
16
/
LETTER #SE82-46.
*17
1
*17
i~1
8
13
A-A
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
Reservoir
Union
Packing
Bolt
Primary Relief Valve
Hydraulic Pump
Packing
Coupling
Self-Relieving Check Valve
Packing
Filter Assembly - Self-Relieving Check Valve
Retainer
Back-Up Ring
Packing
Dipstick
Fluid Filter Screen
Pressure Switch
Packing
Body Assembly
Packing
Packing
Packing
Thermal Relief Valve
Bolt
Washer
Packing
Flat Washer
1
Tee 1
Figure 5A-3.
A
4
/
24
24
26
25
2
NOTE
Assemble new packings and back-up rings
lubricated with a film of Petrolatum W- P236, hydraulic fluid MIL-H-5606, or DowCorning DC-7.
Hydraulic Power Pack Assembly (Sheet 1 of 2)
Revision 3
5A-13
MODEL 210 & T210 SERIES SERVICE MANUAL
-RETURN
VENT
HAND PUMP
SUCTION
Figure 5A-3.
|
Hydraulic Power Pack Assembly (Sheet 2 of 2)
b. Installmotor assembly on top of body assembly (19)
after aligning coupling (8) to match mating connection in
motor. Secure motor to body with bolts. Safety wire bolts
as shown in View A-A.
c. Using new packing (18), install and tighten pressure
switch (17) onto body assembly (19).
d. Using new back-up ring (13), and packings (14) and
(10), install and tighten check valve (9), filter assembly
(11), and retainer (12) into body assembly (19).
e. Install primary relief valve (5) and thermal relief
valve (23) assemblies along with packings (7) and (21)
onto body assembly (19).
CAUTION
Ensure that relief valves are installed in
their correct location. Refer to view A-A.
5A-14
Revision 3
5-1
f. Install fluid filter screen (16) and dipstick (15) into
body assembly (19).
NOTE
Safety wire primary relief valve (5) and
thermal relief valve (23) to hydraulic pump
mounting bolts (4) as shown in view A-A.
g. Using new packings (20) and (26) and washers (25)
and (27), install bolts (24), and tighten reservoir (1) onto
body assembly (19).
5A-10. INSTALLATION. (Refer to figure 5A-3.)
a. Work power pack and bracket assembly into
position and install three bolts, securing power pack to
pedestal.
MODEL 210 & T210 SERIES SERVICE MANUAL
b. Connect all hydraulic lines to power pack fittings.
Ensure that all fittings are properly installed, with
| jamnuts tight, after lines are tightened.
c. Install wheel and gear box assembly and indicator
assembly in top of pedestal.
I d. Install left and right chain guards for rudder trim
chain.
e. Connect chain at connecting link after stringing
chain over idler sprocket.
f. Tighten idler sprocket assembly by sliding sprocket
outboard in slot and tightening bolt.
| g. Connect ground wire to pressure switch (17), and
wire to motor.
h. Connect power pack wiring to plug.
i. Install upper panel assembly on pedestal.
j. Fill reservoir (1) on right side of power pack with
clean hydraulic fluid in accordance with procedures
outlined in Section 2 of this manual.
k. Operate gear through several cycles to bleed
system. Check for correct operation and signs of fluid
leakage. A 28 volt power supply should be used to
augment the ship's battery.
5A-11. PRIMARY AND THERMAL RELIEF VALVE
ASSEMBLIES. (REFER TO FIGURE 5A-3.) The
primary relief valve (5), located between the check valve
(9) and pump (6), serves to limit that amount of pressure
which can be generated by the pump (6). The thermal
relief valve (23), located on the system side of the check
valve (9), serves to limit the system pressure. System
pressure can increase due to thermal expansion. Both
valves are identical except for differing pressure relief
settings (refer to figure 5A-4).
5A-11A. BENCHCHECK OFPRIMARYAND
THERMAL RELIEF VALVES. (Refer to figure 5A-4.)
NOTE
The hydraulic power pack primary relief
valve, thermal relief valve, and pressure
switch can be operationally checked on the
aircraft without power pack removal from the
aircraft or disassembly. Refer to paragraph
5A-5A for specific instructions.
If on-aircraft pressure checking of the power pack
reveals out-of-tolerance relief valve opening, it may be
necessary to determine if relief valve disassembly or
adjustment is necessary. Once removed from power pack,
individual relief valves can be bench checked.
NOTE
Adequate precautions should be taken to
recover hydraulic fluid which will be expelled
from the primary relief valve while under
pressure.
Primary Relief Valve.
(1) Using a hydraulic pump with a flow rate of 0.5 to
0.7 gallons per minute connected to a hydraulic
reservoir, a pressure gage with 2500 psi capacity, and
a hose with appropriate fittings, connect hydraulic
pump to adapter (2) of the primary relief valve.
(2) Apply pressure slowly to ensure that relief valve
assembly opens at correct pressure reading. Primary
relief valve should open at 1800 PSI, + 0 or -50 PSL
Refer to paragraph 5A-15 for adjustment instructions.
b. Thermal Relief Valve.
(1) Using a hand pump connected to a hydraulic
reservoir, a pressure gage with 2500 PSI capacity, and
a hose with appropriate fittings, connect hand pump to
adapter (2) of the thermal relief valve.
(2) Manually pump pressure up slowly to ensure
that relief valve assembly opens at correct pressure
reading. Thermal relief valve should open at 2200
PSI, -0 or + 50 PSI. Refer to paragraph 5A-15 for
adjustment instructions.
a
5A-12. REMOVAL. (Refer to figure 5A-3.)
a. Cut safety wire and remove primary relief valve (5) I
and thermal relief valve (23) from body assembly (19).
5A-13. DISASSEMBLY. (Refer to figure 5A-4.)
NOTE
Relief valve assemblies (5) and (23) are preset
by the factory and normally will not require
disassembly.
a. Removejamnut (13) and adjustmentscrew (12) from |
housing (8).
b. Remove spring (11), guide (10), balls (6), and piston
(9) from housing (8).
c. Loosen jamnut (7) and remove adapter (2) from
housing (8).
d. Remove poppet (4) and orifice (3) from adapter (2).
5A-14. INSPECTION. (Refer to figure 5A-4.)
a. Wash all parts in cleaning solvent (Federal
Specification P-S-611 or equivalent) and dry with filtered
air.
b. Inspect all threaded surfaces for serviceable
condition and cleanliness.
c. Inspect all parts for scratches, scores, chips, cracks,
and indications of excessive wear.
5A-15. ASSEMBLY AND ADJUSTMENT. (Refer to
figure 5A-4.)
NOTE
Use new packings during reassembly.
Lubricate all packings with MIL-H-5606
hydraulic fluid. Lubricate threads with
Petrolatum.
CAUTION
As primary and thermal relief valves are
identical except for differing pressure relief
settings, special care should be exercised to
ensure relief valves are reinstalled in their
correct locations. (Refer to figure 5A-3, view
A-A.)
a. Install orifice (3) and poppet (4) into adapter (2).
(New packing [51 must be installed on poppet [4].)
b. Install jamnut (7) and housing (8) on adapter (2).
c. Tighten adapter (2) into housing (8) and torque to
100-150 lb-in.
Revision 3
5A-15
MODEL 210 & T210 SERIES SERVICE MANUAL
Torque adapter (2) to housing,
(8), and jamnuts (7) and (13)
to housing (8) to 100-150 lb-in.
5A-16. INSTALLATION. (Refer to figure 5A-3.)
a. Install relief valve assemblies (5) and (23) along
with new packings (7) and (21) onto body assembly (19).
CAUTION
MODEL 210 & T210 SERIES SERVICE MANUAL
5A-23. ADJUSTMENT. (Refer to figure 5A-5.)
a. Jack aircraft in accordance with procedure outlined in Section 2 of this manual.
b. Screw cap and housing (10) assembly on fitting
(2) enough to bottom piston out in stop (7).
c. Turn cap and housing assembly (10) back
from full-thread engagement one turn, plus 0,
minus one-fourth turn to locate hole in fitting
(2) in slot in skirt of cap and housing assembly.
d. Attach electrical connections to pressure switch
and attach external power source.
e. Turn on master switch.
f. Pump hand pump to obtain 1500 psi on test gage.
g. Switch should open electrical circuit to pump
solenoid when pressure in system increases to approximately 1500 psi.
h. If switch opens electrical circuit to solenoid
prematurely, disassemble pressure switch down to
washers (8) and add shims as necessary to obtain
desired pressure; repeat steps (b) and (c).
NOTE
Chart in figure 5A-5 lists washers by part
number, thickness and effect in psi each
washer will have on switch operation.
I.
If switch opens electrical circuit to solenoid
later than 1500 * 50 psi, disassemble pressure
switch down to washers (8) and remove washers as
necessary to obtain desired pressure; repeat steps
e. Unscrew guide (5) from fitting (2), do not
damage lip of guide.
5A-24C. INSPECTION AND REPAIR. (See figure
5A-5A. )
a. Clean sealant from threads of snubber (1).
fitting (2) and guide (5) with wire brush.
b. Clean all parts with cleaning solvent (Federal
Specification P-S-661. or equivalent and dry thoroughly.
c. Discard seal (3) and packing (4), and replace
with new parts.
d. Inspect all pressure switch parts for scratches.
scores, chips, cracks and indications of wear.
e. All damaged parts shall be replaced with new
parts.
NOTE
Thorough cleaning is important. Dirt and
chips are the greatest single cause of malfunctions in hydraulic systems. Carefulness
and proper handling of parts to prevent
damage must be observed at all times.
f. Snubber (1) can be cleaned with solvent, then
blown out with high pressure compressed air.
g. Assure that 0. 062-inch vent hole is open in
stop (7).
5A-24D.
REASSEMBLY.
(See figure 5A-5A.)
(b) and (c)..
J.
Turn off master switch.
k.
Lower aircraft to ground.
5A-24. INSTALLATION. (Refer to figure 5A-3.)
Since pressure switch will normally be left in
power pack after adjustment, described in the preceding paragraph, all that needs to be accomplished
is to reassemble the center console. This may be
accomplished by installing the upper panel assembly
on the aft face of the pedestal and installing the decorative cover as outlined in Section 9 of this manual.
5A-24A. REMOVAL AND INSTALLATION.
(21063965 thru 21063972 & 21063974 & ON.)
(See figure 5A-3.)
a. Move left seat to full aft position and spread a
drip cloth beneath power pack.
b. Assure that master switch is OFF, and disconnect leads at terminals at pressure switch.
c. Remove pressure switch from power pack.
d. Reverse procedures for installation.
5A-24B. DISASSEMBLY. (See figure 5A-5A.)
a. Remove pin (10).
b. Unscrew housing (11) from fitting (2).
c. Remove spring (9).
d. Remove washers (8) from flange of stop (7).
NOTE
Chart in figure 5A-5A lists washers by part
number, thickness and effect on operating
pressure (psi).
5A-18
NOTE
Threads of snubber (1) and guide (5) are to
be primed with Locktite grade T primer and
sealed with locktite grade AV sealant. Allow
primer to dry for a minimum of three minutes before sealant application. Allow
sealant to cure from five to 40 minutes
after snubber and guide are assembled.
NOTE
Install new seals and packings and existing
internal parts, lubricated with a film of
Petrolatum W-P-236, hydraulic fluid
MIL-H-5606, or Dow-Corning DC-7. Do
not lubricate threads on guide (5).
a. Install snubber (1) into fitting (2) and tighten with
slotted screwdriver.
b. Install packing (4) in fitting (2).
c. Install seal (3) in guide (5).
d. Install guide (5) into fitting (2), and fingertighten.
NOTE
It is possible to assemble, fill and test the
pressure switch in the aircraft. This can
be accomplished by the installation of a test
gage in the capped port of the tee fitting on
the right-hand side of the power pack, and
pumping the emergency hand pump. Master
switch must be OFF and selector handle
must be in DOWN position.
MODEL 210 & T210 SERIES SERVICE MANUAL
fluid to extend the landing gear in the event of normal
hydraulic pump failure.
5A-27. REMOVAL AND INSTALLATION.
a. Remove seats as required for access.
b. Remove screws attaching cover over hand pump
and remove cover.
c. Peel back carpet as required for access to
pump mounting bolts.
d. Wedge cloth under hydraulic fittings to absorb
fluid, then disconnect the two hydraulic lines and
plug or cap open fittings to prevent entry of foreign
material.
e. Remove two bolts, washers and nuts securing
pump to mounting bracket.
f. Work pump from aircraft.
g. Install hand pump by reversing the preceding
steps, bleeding lines and pump as lines are connected.
h. Fill reservoir as required.
5A-28.
DISASSEMBLY.
(Refer to figure 5A-6.)
NOTE
After emergency hand pump has been removed
from aircraft, and ports are capped or plugged,
spray with cleaning solvent (Federal Specification P-S-611, or equivalent) to remove all accumulated dust or dirt. Dry with filtered compressed air.
a.
Remove hand pump handle by removing pivot and
linkage pins after removing cotter pins.
b. Remove fitting (10) frompump body (16).
c. Push piston (15) from pump body (16).
d. Remove back-up ring (7) from fitting (10) to remove
check valve (8) and KEP-O-SEAL valve (14) assemblies.
e. Remove and discard all O-rings and back-up rings.
5A-18B
Revision 3
5A-29. INSPECTION AND REPAIR
a. Inspect seating surfaces of valves.
b. Inspect piston for scores, burrs or scratches
which could cut O-rings. This is a major cause of
external and internal leakage. The piston may be
polished with extremely fine emery paper. Never
use paper coarser than No. 600 to remove scratches
or burrs. If defects do not polish out, replace
piston.
5A-30. REASSEMBLY. (Refer to figure 5A-6).
Assemble the emergency hand pump, using the figure
as a guide. Also, for detailed instructions, reverse
the procedures outlined in paragraph 5A-28. During
assembly, prime fitting (10) with Locktite grade T
primer, allow primer to dry for a minimum of three
minutes. Apply Locktite hydraulic sealant to threads
of pump body (16) and first two threads of the fitting
(10). After installing fitting in pump body, allow the
sealant to cure from five to 40 minutes.
NOTE
Install new back-up rings and packings,
lubricated with a film of Petrolatum
VV-P-236,
hydraulic fluid MIL-H-5606,
or Dow-Corning DC-7.
5A-31. LANDINGGEAR SELECTOR VALVE.
to figure 5A-7. )
(Refer
5A-32. DESCRIPTION. A mechanical gear position
selector valve is located in the switch panel The
pilot shuttles the valve mechanically when he changes
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Beginning with 21063812, steel disc (9) is
replaced by aluminum disc (9). Bearing
(5) and race (4) are replaced by teflon
washer (17) at the same aircraft serial.
10
.
1.
2.
3.
4.
5.
6.
7.
8.
Cover
Retaining Ring
Cap
Bearing Race
Thrust Bearing
Washer
Spring
Packing
9.
10.
11.
12.
13.
14.
15.
16.
17.
16
.7
9
Disc
Packing
Pucks
Spring
Body Assembly
Rod
Nut
Knob
Washer
Figure 5A-7.
5A-35. INSPECTION AND REPAIR (Refer to figure
5A-7.) Replace packings (8) and (10). Check valve
for wear, foreign or abrasive materials. Disc (9)
may be refaced (lapped) if worn or abraded. Check
rollers in bearings (5).
5A-36. RIGGING THROTTLE-OPERATED GEAR
WARNING BORN MICRO-SWITCH. (Refer to figure
5A-8.) Rigging procedures for sea level or turbocharged aircraft are outlined in figure A-8. )
5A-20
2
1
Landing Gear Position Selector Valve
to figure 5A-7.)
a. Remove cover (1), retaining ring (2) and cap (3).
Thru 21063811, remove race (4) and bearing (5). Beginning with 21063812, remove washer (17).
b. Remove cotter pin, washer (6) and spring (7).
c. Pull rod (14) from disc (9); remove disc.
d. Remove packs (11) and springs (12).
e. Reverse preceding steps for reassembly.
5A-37.
3
MAIN LANDING GEAR. (Refer to figure
5A-9.)
5A-38. DESCRIPTION. The tubular main gear
struts rotate aft and inboard to stow the main wheels
beneath the baggage compartment. The main gear
utilizes hydraulic pressure for positive uplock and
mechanical downlocks. Main gear uplock pressure
is maintained automatically by the pump assembly.
Rotation of the gear to extend or retract the struts
is achieved through pivot assemblies which in turn
are bolted through a splined shaft, to the hydraulic
main gear rotary actuators.
TCuTIa
Use of recapped tires or new tires not
listed on the aircraft equipment list are
not recommended due to possible interference between the tire and structure
when landing gear is in the retracted
position.
MODEL 210 & T210 SERIES SERVICE MANUAL
SETTING THROTTLE SWITCHES
1. During flight at 120 MPH (IAS), 2500', prop
control full forward for maximum RPM, and with
the gear and flaps up, mark the throttle control
position corresponding to the following manifold
pressures:
12. 0" ± .5"
15.0"+ 1.0"
Model 210M
Model T210M
2. Then adjust the gear warning horn throttle
switch on the ground to activate at the throttle
control position as marked in flight.
"For each 1000 feet above 2500' MSL, decrease
the manifold pressure at which the throttle con-
E
TO
SECTION 2
FOR CONTROL
LUBRICATION
6
Figure 5A-8.
1.
2.
2.
3.
3.
4.
5.
5.
Rigging Throttle-Operated Gear Warning Horn Switch
VIEW LOOKING AFT AND
RIGHTOUTBOARD AT RIGHTHAND SIDE OF FIREWALL
Cover
Switch Cover
Switch
Spacer Cover
Spacer
Switch
Switch
Switch Spacer
Figure 5A-8.
6.
7.
7.
8.
8.
9.
10.
5A-21
Switch Mounting Bracket
Arm
Warning Cam
Gear Assembly
Gear Warning Cam
Fuel Pump Cam
Bushing
Rigging Throttle-Operated Gear Warning Horn Switch
5A-21
MODEL 210 & T210 SERIES SERVICE MANUAL
5A-39.
TROUBLE SHOOTING.
TROUBLE
AIRCRAFT LEANS TO
ONE SIDE.
UNEVEN OR EXCESSIVE
TIRE WEAR.
PROBABLE CAUSE
REMEDY
Incorrect tire inflation.
Inflate to correct pressure.
Sprung main gear strut.
Remove and replace strut.
Bent axle.
Install new axle.
Incorrect tire inflation.
Inflate to correct pressure.
Wheel out of alignment.
Align wheels.
Wheels out of balance.
Balance wheels.
Sprung main gear strut.
Replace strut.
Bent axle.
Install new axle.
Dragging brakes.
Jack wheel and check brake.
Wheel bearings not adjusted
properly.
Tighten axle nut properly.
5A-40. REMOVAL. (Refer to figure 5A-9.)
a. Jack aircraft in accordance with procedures
outlined in Section 2 of this manual.
b. Bleed fluid from brake line at wheel brake
cylinder.
c. Turn master switch off; move gear position
selector valve to up position, then turn master switch
Place gear position selector handle in a neutral position so that gear rotates freely.
f. Remove packings (24) from plug (25) and clean plug |
and strut (29).
5A-41.
INSTALLATION. (Refer to figure 5A-9.)
NOTE
The following procedure installs the landing
NOTE
If the pump motor cannot be used to unlock
the main gear because of an opening in the
hydraulic system, the spring-loaded main
gear downlocks can be manually unlocked
a. Lubricate new -rings (24) and end of strut (29)
with Petrolatum W-P-236, hydraulic fluid MIL-L-5606,
or Coming DC-7 (keep DC-7 away from areas to be
painted) before installation Install O-rings (24) on plug
(25).
b. Remove caps from union (23) and brake line (22),
attach brake line (22) to union (23), and work plug (25)
by pushing them forward with a screw-
and strut (29) into pivot (14).
driver or other similar tool, and holding
them forward, until the main gear has rotated past.
NOTE
When installing a new pivot (14), burnishing
WARNING
the 2.100-inch I. D. bore may be required to
It is advisable to have an assistant hold the
gear strut up while the locks are pushed
forward to prevent the strut from rotating
suddenly, possibly causing personal injury.
Ensure that master switch is OFF and pump
motor circuit breaker pulled.
c. Align hole in plug (25) with holes in pivot assembly
(14) using special tool No. SE934.
d. Remove strut attach bolt (26) and work strut (29)
and plug (25) from pivot assembly (14).
e. Disconnect brake line from union (23) and plug
union and brake line.
5A-22
Revision 3
facilitate assembly of landing gear strut (29).
NOTE
Special tool No. SE934 is available from
Cessna Parts Distribution CPD 2) through
Cessna Service Stations. This tool is designed
to install strut attaching bolt (26) without
damaging the packings (24) in the plug (25).
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Tie
Boot
Pin
Spacer
Arm Assembly
Downlock Adjustment Screw
Hook
Setscrew
Down LimitSwitch
Spring Assembly
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
Eyebolt
Shim
Support Assembly
Landing Gear Strut
Shell Assembly
Stud
Pivot Assembly
Bolt
Rod End
Actuator
Figure 5A-10.
5A -24
Revision 3
NOTE
Stud (16) and downlock support shell (15)
attaching screws shall be sealed with
grade AV Locktite 271, or Locktite Catalog #87. Allow sealant to cure for 12
hours before service use.
Rigging Main Landing Gear (Sheet 1 of 3)
MODEL 210 & T210 SERIES SERVICE MANUAL
Downlock Hook Overcenter
Downlock Hook Overcenter
Downlock Hook Overcenter Gage (SE960)
MS 20392 PIN
Setscrew
Gage (SE960)
Downlock
Actuator RodHook
Shoulder
Spacer
ASSEMBLIES
RIGGING DOWNLOCK
Arm Assembly
Limit
Pivot
Assembly
WHEN
LOCK IS
DOWNLOCK HOOK OVERCENTERNESS
MORE THAN MINIMUM TOLERANCE
ASSEMBLIES
RIGGING DOWNLOCK
OUTBOARD
LOOKING
Downlock Hook
MS 20392 PIN
Downlock
Actuator Rod
Hook
Shoulder
Arm Assembly
Assembly Switch
DOWNLOCK HOOK OVERCENTERNESS
LESS THAN MAXIMUM TOLERANCE
Downlock Hook Overcenter Gage (SE960)
Down
Spacer
0. 18± . 02
WHEN
LOCK IS
LOCKED
pivot
Assembly
ABBREVIATIONS ON GAGES:
NHLT _ NOT HITTING, LESS THAN DIMENSION STAMPED
ON GAGE
NHMT _ NOT HITTING, MORE THAN DIMENSION STAMPED
ON GAGE
Figure 5A-10.
Rigging Main Landing Gear (Sheet 2 of 3)
Revision 3
5A-25
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
with a preassembled length of 12.45-inches, and the three
hydraulic ports in the same plane. Install actuator
assembly (20), attaching it to fuselage structure and arm
assembly (5).
d. With landing gear free, hydraulic pressure off,
and downlock systems in position shown on sheet 1,
swing landing gear into the DOWN position and adjust
adjusting screw as follows:
NOTE
To relieve hydraulic pressure, pull hydraulic
pump circuit breaker off, and move gear selector handle up and down two or three times.
1. If downlock locks, turn adjusting screw 1/4
turn out at a time until lock will not lock; then turn
back in 1/4 turn and secure pin.
2. If downlock does not lock, turn adjusting
screw 1/4 turn in at a time until lock will lock, then
secure pin.
| e. Readjust setscrew (8) to stop hook assembly .040 to
.090-inch overcenter. When checking overcenter
measurement of arm assembly (5), landing gear should
be as shown on sheet 2, with nut, washer, and spacer
removed, which retains the arm assembly (5). Use
downlock overcenter gages (P/N SE960) to determine if
hook (7) is still within tolerances shown on sheet 2. Use
gages as follows:
6. If a slight rotation is possible, setscrew (8) is not
contacting downlock adjustment screw (6). If contact is
not being made, downlock actuator will have to be
readjusted by backing offactuator's rod end (19) one-half |
turn at a time (one-and-one-half turns maximum
adjustment) until hook assembly is 0.040-inch or more
overcenter, and contact is being made between setscrew
(8) and downlock adjustment screw. If contact is being
made, setscrew (8) should be adjusted outward to increase
overcenter measurement to within tolerance.
NOTE
For correct rigging, downlock hook setscrew
(8) must make contact with downlock
adjustment screw (6) and green areas of both
gages must contact as shown on sheet 2.
f. Now that downlock adjustment screw (6) has been
I
adjusted following procedures outlined in step "e.", check,
downlock actuator rod end (19) adjustment as follows:
1. Connect all hydraulic lines, fill system with
MIL-H-5606 hydraulic fluid and purge system of air by
cycling gear through several cycles.
NOTE
Check fluid level in power pack reservoir
frequently during purging and rigging procedures.
NOTE
Overcenter gages (P/N SE960) are available
from Cessna Parts Distribution (CPD 2)
through Cessna Service Stations.
1. Remove nut, washer and spacer which retain
arm assembly to support assembly.
2. Install 0. 090-inch downlock gage (SE960) on
inboard side of downlock hook as shown on sheet 2.
Upper portion of gage should rest against head of pin
attaching adjusting screw. If downlock hook is under
maximum overcenter tolerance, green area of gage
will contact spacer on gear pivot, while red area will
not make contact with 0. 50-inch diameter shoulder,
as shown in the figure. When downlock hook is in
maximum overcenter tolerance, both green and red
areas will make contact. If red area makes contact
and green area does not, the downlock hook setscrew
should be adjusted INWARD to bring overcenter dinension within tolerance.
3. Install 0. 040-inch downlock gage (SE960) on
inboard side of downlock hook as shown on sheet 2.
If downlock hook is over minimum overcenter tolerance, green area of gage will contact shoulder, while
red area will not make contact with spacer.
4. When downlock hook is in minimum overcenter tolerance both red and green areas will
make contact
1
ipivot
5. If overcenter tolerance is less than 0.040-inch,
the red area will make contact, while the green area will
not. If this condition exists, the next step is to determine
if the downlock adjustment screw (6) is making contact
with the setscrew (8). This is accomplished by lifting the
landing gear spring upward off the hook assembly and
checking for possible rotation of the hook assembly, by
hand, with hydraulic pressure off.
2. Pull hydraulic pump circuit breaker off.
3. With gear in down and locked position, move
gear selector handle to GEAR UP position and note
actuation of main gear downlock hooks.
4. As soon as left downlock hook is actuated to
unlock the left gear, move gear selector handle back
to GEAR DOWN position to simulate what would occur
if the pilot were to select gear down before the gear
was fully retracted. If downlock hooks do not lock
the gear in the down position, check downlock system
for misalignment.
5A-43 RIGGING MAIN GEAR DOWN LIMIT
SWITCHES. (Refer to figure 5A-10, sheets 1 and 2.) The
main gear down limit switches (9) are attached to
brackets which are welded to the spring assembly (10).
Adjustment is accomplished by loosening the lock nut
and either tightening or loosening the adjustment nut
and retightening the lock nut against the bracket behind
the adjustment nut. Down limit switches (9) are to be
adjusted to the dimension stipulated in Sheet 2.
5A-44. RIGGING MAIN GEAR UP LIMIT SWITCHES.
(Refer to figure 5A-10, Sheet 3.) The main gear up limit
switches (6) are mounted in indicator light brackets (2)
which are attached to the underside of the removable
floorboards (1), immediately above the main landing gear
assemblies. The switches are contacted by
actuators, bonded to clamps, which are attached to
the aft leg of the landing gear strut pivot assembly.
When replacing a clamp/actuator assembly, adjust
the actuator tab prior to bonding, so that it actuates
the gear-up indicator light switch. Bond the actuator
to the clamp with HYSOL EA-9309 or 3M EC-2216
Revision 3
5A-27
MODEL 210 & T210 SERIES SERVICE MANUAL
PLACE CARPENTER'S SQUARE
AGAINST STRAIGHTEDGE AND
LET IT TOUCH WHEEL JUST
MODEL 210 & T210 SERIES SERVICE MANUAL
by unscrewing from actuator body (3).
b. Remove cap (1) from end of actuator,
c. Using a small rod, push piston (12) from actuator body.
NOTE
Unless defective, do not remove nameplate.
bearings (2) or roller (13).
d. Remove packing (5) and back-up ring (4) from
cylinder body (3). Discard packing (10).
e. Remove packing (10) and back-up ring (9) from
end gland (8). Discard packing (10).
f. Remove and discard packing (11) from piston
(12).
5A-52. INSPECTION.
a. Thoroughly clean all parts in cleaning solvent
(Federal Specification PS-661, or equivalent. )
b. Inspect all threaded surfaces for cleanliness,
cracks and wear.
c. Inspect cap (1), piston (12), roller (13), if removed, and actuator body (3) for cracks, chips,
scratches, scoring, wear or surface irregularities
which may affect their function or the overall operation of the actuator.
d. Inspect bearings (2), if removed, for freedom
of motion, scores, scratches or Brinnel marks.
5A-53.
PARTS REPAIR/REPLACEMENT.
Repair
of small parts of the main landing gear actuator is
impractical. Replace all defective parts. Minor
scratches or score marks may be removed by polishing with abrasive crocus cloth (Federal Specification
P-C-458), providing their removal does not affect
operation of the unit. During assembly, install all
new packings.
5A-54.
REASSEMBLY.
(Refer to figure 5A-12.)
NOTE
Install new packings and back-up rings
lubricated with a film of Petrolatum
VV-P-236, hydraulic fluid MIL-H-5606,
or Dow-Corning DC-7. If roller (13)
and bearings (2) have been removed,
lubricate with MIL-G-2116C lubricant.
a. If bearings (2) and roller (13) were removed,
press one bearing into actuator body until it is flush.
Install roller and press second bearing in place to
hold roller. Use care to prevent damage to bearings
or roller.
b. Install back-up ring (4) and packing (5) in actuator body core. Install new packing (11) and back-up
rings (6) on piston (12).
NOTE
Lubricate piston rack gears with MIL-G21164C lubricant. Apply lubricant sparingly. Over-greasing might cause contamination of hydraulic cylinder assembly
with grease which might work past packing.
c.
d.
end
e.
Slide piston (12) into cylinder body (3).
Install back-up ring (9) and new packing (10) on
gland.
Install end gland in cylinder and tighten until end
of gland is flush with end of cylinder body. Install
and tighten setscrew (8).
f. Install cap (1) at end of actuator assembly.
5A-55. INSTALLATION.
a. With main landing gear in the down and locked
position, install actuator into bulkhead forging so
that piston rack gear and sector gear engage as
shown in figure 5A-9, Section A-A.
b. Lubricate swivel fitting on actuator with MIL-G21164 lubricant, install packing in fitting.
c. Install cap (4), washer (3), retainer (2) and
swivel fitting on actuator as shown in figure 5A-9.
d. Install bolts (-and torque to 60-85 lb in. Safety
wire swivel fitting to shaft (8).
e. Connect all hydraulic lines to their source locations. Lubricate threads with Petrolatum, W-P236.
f. Connect brake line at wheel cylinder. Fill and
bleed brake system in accordance with procedures
outlined in applicable paragraph in this section.
g. Rig landing gear in accordance with procedures
outlined in applicable paragraph in this section.
h. Remove aircraft from jacks and Install access
covers, carpeting and seats removed for access.
5A-56
MAIN GEAR PIVOTASSEMBLY.
5A-57.
REMOVAL. (Refer to figure 5A-9.)
5a.
Remove strut from pivot assembly in accordance
with procedures outlined in applicable paragraph in
with procedures outlined in applicable paragraph in
this section
b. Remove actuator in accordance with procedures
outlined in applicable paragraph in this section.
c. Remove setscrew from sector gear (7).
d. Bend tangs of washer (21) from notches in nut
(20) and completely unscrew nut (20) from threaded
area of shaft (17).
e. Push shaft (17) into pivot assembly (14) and pull
pivot assembly free of shaft (8).
5A-58. INSPECTIONAND REPAIR. (Refer to figure
5A-9. )
a. Thoroughly clean all parts in cleaning solvent
(Federal Specification PS-661 or equivalent. )
b. Inspect all parts for indications of damage,
cracks or excessive wear and replace as necessary.
c. Inspect outboard pivot bushing and inboard pivot
bearing (10) (pressed into bulkhead forgings in aircraft) for damage and excessive wear. Replace bushing or bearing as required.
NOTE
The outboard pivot bushing is locked into the
bulkhead forging by a setscrew located above
the bushing. This setscrew must be turned
out several turns before the bushing can be
removed.
5A-59. INSTALLATION. (Refer to figure 5A-9.
a. Lubricate all bushings and bearings with MIL-G21164 grease. Slide shaft (17) into pivot assembly
(14).
b. Install pivot with bearing (12) and race (11) installed, into inboard bearing in bulkhead forging.
Pull shaft from pivot and install washer (211 and nut
(20) on shaft.
5A-31
MODEL 210 & T210 SERIES SERVICE MANUAL
c. Insert end of shaft into outboard bushing in bulkhead forging. Hand-tighten nut to remove all end
play and safety in place by bending corresponding
tang of washer into notch of nut. Pivot must rotate
freely.
d. Install seal (9) and sector gear (7) on inboard
end of pivot assembly so that setscrew hole in sector
gear lines up with setscrew hole in shaft (8); install
setscrew into sector gear and shaft with Loctite 242
locking compound and tighten screw.
5A-60.
GEAR POSITION INDICATOR SWITCHES.
5A-61. DESCRIPTION. The gear down indicator
switches are attached to brackets which are welded
to the downlock hooks. The main gear up limit
switches are mounted in brackets which are attached
to the underside of the removable floorboards immediately above the main landing gear pivot assemblies. Refer to the paragraphs in this section which
outline procedures for rigging the main gear up and
down switches.
5A-62. MAIN GEAR DOWNLOCK ACTUATOR.
(Refer to Section 5.)
b. Mark position of removed step so new step will
be installed in approximately the same position on
the strut.
c. Check that bonding surfaces are clean and
thoroughly dry.
d. Mix adhesive (Uralite 3121 or 3M EC-2216 per
manufacturer's instructions. Note pot life.
e. Spread a coat of mixed adhesive on bonding surfaces of strut and step; install step on strut.
NOTE
Top of strut should be parallel to the ground
(±5°) when gear is in down position.
I. Cycle landing gear to check clearance of step in
tunnel.
g. Form a small fillet of adhesive at all edges of
bonding surfaces. Remove excess adhesive.
h. Remove aircraft from jacks.
i. Allow adhesive to thoroughly cure according to
manufacturer's recommendations before flexing gear
spring or applying loads to step.
j. Paint gear spring and step after curing is completed.
5A-68.
5A-63. DESCRIPTION. The main gear downlock
actuators for the 1979 Models is the same actuator
used on Models thru 1978. Function and operation
are the same. The only difference between the actuators is the replacement of the MS28778-4 fitting with
a hose assembly. Refer to Section 5 for actuator remuval, disassembly, inspection and repair and installation. Adjustment of the actuator rod end is
discussed in the main landing gear rigging paragraph
in Section 5A.
5A-64.
56A-9.)
MAIN GEAR STRUT STEP.
(Refer to figure
5A-65. DESCRIPTION. The step is constructed of
Uralite 3121 polyurethane costing, with a molded
depression area, located in the top of the step. An
adhesive-backed "Walkway" material with rough surface Is pressed into the depressed area of the strut.
5A-66.
REMOVAL.
NOTE
Step is bonded to gear spring with Uralite
3121 or 3M EC-2216 adhesive.
.. Using
a heat gun, heat step at a temperature of
NOSE GEAR SYSTEM.
5A-69, DESCRIPTION. The nose gear consists of
a pneudraulic shock assembly, mounted in a trunnion assembly, a steering arm and bungee, shimmy
dampener, nose wheel, tire and tube, hub cap, bearing, seals and a double-acting hydraulic actuator for
extension and retraction. A claw-like hook on the
actuator serves as a downlock for the nose gear.
SA-70. OPERATION. The nose gear shock strut is
pivoted just forward of the firewall. Retraction and
extension of the nose gear is accomplished by a
double-acting hydraulic cylinder, the forward end of
which contains the nose gear downlock. Initial action
of the cylinder disengages the downlock before retraction begins. Nose gear doors are mechanically
closed as the nose gear retracts. As the nose gear
extends, the doors are mechanically opened.
5A-71. TROUBLE SHOOTING. Refer to the nose
gear system trouble shooting chart in Section 5.
5A-72. REMOVAL OF NOSE GEAR ASSEMBLY.
Refer to applicable paragraphs in Section 5, outlining nose gear removal, disassembly, inspection
and repair, reassembly and installation, disregarding the installation step regarding rigging of the retractable step.
200º -250- F, until step material becomes pliable.
b. Using a sharp knife, remove step material down
to the metal strut.
c. Clean off remaining step material with a wire
wheel and sandpaper. Leave surface slightly rough
or abraded. Clean oil and grease from strut with
solvent, wipe off excess solvent with dry cloth and
let surface dry.
d. Apply zinc chromate, primer - green or yellow
to cleaned area on struts. Dry film thickness to be
0003 to. 0005 inch.
5A-74. TORQUE LINKS. Refer to applicable paragraphs in Section 5 outlining removal of torque links
and squat switch.
5A-67. INSTALLATION.
a. Jack aircraft in accordance with procedures
outlined in Section 2 of this manual.
5A-75. SQUAT SWITCH. Refer to applicable paragraphs in Section 5 outlining removal and installation
of torque links for squat switch removal.
5A-32
5A-73. SHIMMY DAMPENER Refer to applicable
paragraphs in Section 5 outlining description, removal, disassembly, inspection, repair and reassembly of the shimmy dampener.
MODEL 210 & T210 SERIES SERVICE MANUAL
5A-76.
NOSE GEAR DOWNLOCK MECHANISM.
quired.
Refer to applicable paragraphs in Section 5 outlining
-
description, removal, disassembly, inspection, re-
CAUTION
pair and reassembly of the nose gear actuator.
5A-77. NOSE GEAR ACTUATOR. Refer to applicable paragraphs in Section 5 outlining description,
removal, disassembly, inspection, repair and reassembly of the nose gear actuator.
.5A-78. NOSE GEAR DOOR SYSTEM.
figure 5A-13.)
(Refer to
5A-79. DESCRIPTION. The nose gear door system
consists of a right and left forward door. actuated
by push-pull rods and a torque tube assembly. The
aft doors are attached to the torque tube assembly
with springs.
The piston rod is flattened near the threads
to provide a wrench pad. Do not grip the
piston rod with pliers, as tool marks will
cut the O-ring seal in the actuator.
5A-84. RIGGING NOSE GEAR DOWN LIMIT SWITCH.
(Refer to figure 5A-14.) The nose gear down limit
switch is mounted on a tab which is a part of the bearing end (5)
the nose gear actuator. The switch is actuated by the right-hand actuator locking hook (1)
Switch adjustment is accomplished by loosening the
ment nut and re-tightening the lock nut against the
tab behind the adjustment nut.
Down limit switch is
to be adjusted to the dimension stipulated in the figure.
5A-80. REMOVAL AND INSTALLATION. (Refer to
figure 5A-13. )
a. Remove hinge bolts, nuts, washers and bushings.
b. Remove nuts from push-pull rods and remove
forward doors,
c. Disconnect spring from aft door eyebolt, and
remove aft doors.
d. Reverse preceding steps to install nose gear
doors.
5A-85. RIGGING NOSE GEAR UP LIMIT SWITCH.
(Refer to figure 5A-14.) The nose gear up limit
switch is mounted to a bracket, located in the lefthand forward area of the nose wheel well. The
switch is activated by the left-hand arm of the bellcrank weld assembly. Switch adjustment is provided
by slots in the switch mounting bracket. Up limit
switch is to be adjusted to the dimension stipulated
in the figure.
NOTE
Upon completion of installation, safety wire
bolts (*) to clips (23).
NOTE
Check nose gear door-to-cowling clearance
to be 0.12-inch to 0.15-inch on the left and
right sides of the nose gear doors each time
the turbine access door on turbocharged
models is re-installed.
5A-81.
NOSE WHEEL STEERING SYSTEM.
5A-82. DESCRIPTION. Refer to applicable paragraphs in Section 5, outlining description, removal,
installation and rigging of the nose wheel steering
system,
5A-83. RIGGING NOSE LANDING GEAR. (Refer to
figure 5A-14.)
NOTE
Nose gear shock strut must be correctly
inflated prior to rigging the nose gear.
Refer to Section 1 of this manual for correct nose gear shock strut inflation pressure.
a. Jack aircraft in accordance with procedures
outlines in Section 2 of this manual.
b. Actuator locking hooks (1) on the nose gear
actuator shall completely engage downlock pins (2)
without drag, and cross bar (3) shall rotate freely
to indicate it is not bearing on either side of slot
in rod end (4). Adjust rod end of actuator as re-
5A-86. RIGGING OF NOSE GEAR SQUAT SWITCH.
(Refer to figure 5A-14.) The nose gear squat (safety)
switch is mounted in a bracket, attached to the upper
nose gear torque link. The switch is operated by an
actuator, attached to the nose gear lower torque link.
Adjust squat switch so that contacts close when nose
gear strut is .12 to .25-inch from fully-extended position.
5A-87.
RIGGING OF NOSE GEAR DOORS. (See figure 5A-13.) Nose gear door adjustments are accomplished by adjusting push-pull rod ends as required
to cause the doors to close snugly. Doors must fair
when the nose gear is fully retracted. Link rods are
to be adjusted so that the doors, when in the open position, clear any part of the nose gear assembly by a
minimum of 0. 25-inch during retraction. Adjust stop
bolts on stop assemblies (12) as required to contact
arms (9) on bellcrank weld assembly (15) when forward
nose gear doors are in FULL-OPEN position. Adjust
barrel assemblies (4) as required to fair forward nose
gear doors in closed position. Pack bearings (16)
with MIL-G-21164 grease. Trim outboard edge of
forward nose gear doors so that door-to-skin clearance is 0.18-inch minimum to 0.21-inch maximum.
Safety wire bolts (*) to clips (23).
5A-88. FINAL LANDING GEAR SYSTEMS CHECK.
After landing gear systems have been installed and
rigged, prior to removal from jacks, cycle landing
gear through 25 cycles using the system's emergency
hand pump.
NOTE
Check fluid level in power pack reservoir
frequently during purging and system checks.
5A-33
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
(Must rotate freely
0.
18± 0. 02-inch when lock is locked
DOWN LIMIT
SWITCH
NOSE GEAR IN DOWNLOCK POSITION
BRACKET
SWITCH
FORK
ROD END
1. 05-inch when gear fork is against stop
SPRING
NOSE GEAR IN UP POSITION
Figure 5A-14.
5A-36
Revision 3
Rigging Nose Landing Gear (Sheet 2 of 2)
NOSE GEAR
ACTUATOR
MODEL 210 & T210 SERIES SERVICE MANUAL
One of the 25 cycles shall consist of a downlock malfunction check, consisting of the following procedure,
using a 28 volt DC, 60 amp electrical power supply.
a. Pull hydraulic circuit breaker off.
b. With gear in down and locked position, move
gear selector handle to GEAR UP position and note
actuation of main gear downlock hooks.
c. As soon as left downlock hook is actuated to
unlock the left gear, move gear selector handle back
to GEAR DOWN position to simulate what would occur
if the pilot were to select gear down before the gear
was fully retracted. If downlock hooks do not lock
the gear in the down position, check downlock system
for misalignment.
NOTE
This malfunction check is in addition to the
check used during the rigging procedure.
d.
Remove aircraft from jacks.
5A-89. NOSE WHEEL AND TIRE. Refer to applicable paragraphs in Section 5, outlining description,
removal, disassembly, inspection, repair, reassembly and installation of nose wheels and tires.
5A-90. BRAKE SYSTEM. Refer to applicable paragraphs in Section 5 for description, trouble shooting,
removal, disassembly, inspection, repair, reassemgly, installation, checking lining wear, lining installation and bleeding of the brake system. Refer to the
following note.
NOTE
Approximately 200 of the initial 1979 production model aircraft may be equipped
with brake assemblies having 1/4-inch
fittings in lieu of 3/16-inch fittings. Refer
to the Model 210 Parts Catalog for replacement parts.
5A-91. BRAKE MASTER CYLINDER.
figure 5A-15. )
(Refer to
5A-92. DESCRIPTION. The brake master cylinders,
located immediately forward of the pilot's rudder
pedals, are actuated by applying pressure at the top
of the rudder pedals. A small reservoir is incorporated into each master cylinder for the fluid supply.
When dual brakes are installed, mechanical linkage
permits the copilot pedals to operate the master
cylinders.
5A-93. REMOVAL.
a. Remove bleeder screw at wheel brake assembly
and drain hydraulic fluid from brake cylinders.
b. Remove front seats and rudder bar shield for
access to brake master cylinders.
c. Disconnect parking brake linkage and disconnect
brake master cylinders from rudder pedals.
d. Disconnect hydraulic hose from brake master
cylinders and remove cylinders.
e. Plug or cap hydraulic fittings, hose and lines
to prevent entry of foreign material.
5A-94. DISASSEMBLY. (Refer to figure 5A-15.)
a. Unscrew clevis (1) and jam nut (2).
b. Remove filler plug (3).
c. Unscrew cover (4) and remove up over piston
(5).
d. Remove piston (5) and spring (8).
e. Remove packing (7) and back-up ring (6) from
piston (5).
5A-95. INSPECTION AND REPAIR. (Refer tofigure
5A-15.) Repair is limited to installation of new parts
and cleaning. Use clean hydraulic fluid (MIL-H-5606)
as a lubricant during reassembly of the cylinder. Replace packings and back-up rings. Filler plug (3)
must be vented so pressure cannot build up during
brake operation. Remove plug and drill 1/16-inch
hole, 30 ° from vertical, if plug is not vented. Refer
to view A-A for location of hole.
5A-96. REASSEMBLY. (Refer to figure 5A-15.)
a.Install spring (8) into cylinder body (9).
b. Install back-up ring (6) and packing (7) in groove
of piston (5).
c. Install piston (5) in cylinder body (9).
d. Install cover (4) over piston (5) and screw cover
into cylinder body (9).
e. Install nut (2) and clevis (1).
f. Install filler plug (3), making sure vent hole is
open.
5A-97. INSTALLATION.
a. Connect hydraulic hoses to brake master cylinders.
b. Connect brake master cylinders to rudder pedals
and connect parking brake linkage.
c. Install rudder bar shield and install front seats.
d. Install bleeder screw at wheel brake assembly
and fill and bleed brake system in accordance with
applicable paragraph in Section 5.
5A-96. PARKING BRAKE SYSTEM. Refer to applicable paragraphs in Section 5 for description, removal, installation, and inspection and repair of
components of the parking brake system.
5A -37
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 6
AILERON CONTROL SYSTEM
Page No.
Aerofiche/Manual
TABLE OF CONTENTS
AILERONCONTROL SYSTEM .....
Description ....................
Trouble Shooting
.........
Control Column ................
Description .
............
Removal and Installation .....
Repair ......................
Bearing Roller Adjustment ...
Aileron Bellcrank ..............
Removal
.
.............
Installation .................
6-1. AILERON CONTROL SYSTEM.
ure 6-1.)
6-2.
DESCRIPTION.
6-3.
TROUBLE SHOOTING.
1K16/6-1
1K16/6-1
1K16/6.1
1K17/6.2
1K17/6-2
1K17/6-2
1K24/6-9
1K24/6.9
1K24/6-9
1K24/6-9
1L1/6-10
Repair .................
Ailerons ....................
Removal and Installation
Repair ......................
Aileron Trim Tab ..........
Removal and Installation
Adjustment ............
Cables and Pulleys .........
Removal and Installation
Rigging .....................
(Refer to fig-
.
1L/6-10
1L6-10
1L/6.10
1L6-10
1L6-10
1L16-10
L2/6-11
1L2/611
1L2/6-11
1L2/6.11
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 rerig 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.
Revision 3
6-1
MODEL 210 & T210 SERIES SERVICE MANUAL
6-3.
TROUBLE SHOOTING (Cont).
TROUBLE
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 alignmest
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.
CONTROL WHEELS NOT
LEVEL WITH AILERONS
NEUTRAL.
6-4. CONTROL COLUMN (Refer to figure 6-2.)
6-5: DESCRIPTION. (Refer to figure 6-2, Sheets 1
and 2.) Rotation of the control wheel rotates four
bearing roller assemblies (15) on the end of the control
wheel tube (14), which in turn, rotates a square control
tube assembly (20) inside and extending from the control
wheel tube (14). Attached to this square control tube
assembly (20) is a quadrant (29) 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 (32),
interconnect cable turnbuckle (33), and interconnect
cable adjustment terminals (28). The forward end of the
square control tube assembly (20) is mounted in a
bearing block (31) on firewall (34) and does not move
fore-and-aft, but rotates with the control wheel. The
four bearing roller assemblies (15) on the end of the
control wheel tube (14) reduce friction as the control
wheel is moved fore-and-aft for elevator system
operation. A sleeve weld assembly (11), containing
bearings which permit the control wheel tube (14) to
rotate within it, is secured to the control wheel tube (14)
by a sleeve and retaining ring in such a manner that it.".
moves fore-and-aft with the control wheel tube. This
movement allows the push-pull tube (22), attached to the
sleeve weld assembly (11), to operate an elevator arm
assembly (23), to which one elevator control cable (24) is
attached. A torque tube (37) connects this elevator arm
assembly (23) to the one on the opposite end of the torque
tube (37), 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. (Refer to figure 6-2, Sheet 3.) Slide cover (2)
toward instrument panel to expose adapter (3). Remove
bolts securing adapter (3) to control wheel tube (1).
6-2
Revision 3
b. Disconnect electrical wiring to map light, mike
switch, and electric trim switch at connector (4), if
installed. Slide cover (2) off control wheel tube (1).
c. (Refer to figure 6-2, Sheets 1 and 2.) Remove
decorative cover from instrument panel.
d. Remove screw securing glide plug (18) to control
tube assembly (20) and remove glide plug (18) and glide
(19).
e. Disconnect push-pull tube (22) at sleeve weld
assembly (11).
f. Remove screws securing cover plate (5) at
instrument panel.
g. Using care, pull control wheel tube (14) aft and
work assembly out through instrument panel.
NOTE
To ease removal of control wheel tube (14),
snap rings (7) may be removed from their
locking grooves to allow sleeve weld assembly (11) additional movement.
If removal of control tube assembly (20) or
quadrant (29) is necessary, proceed to step
h. Remove safety wire and relieve direct cable tension
at turnbuckles (Index 5, figure 6-1).
i. Remove safety wire, relieve interconnect cable
turnbuckle (33) tension, and remove cables from
quadrant (29).
j. Remove safety wire and remove roll pin (25)
through quadrant (29) and control tube assembly (20).
k. Remove pin, nut (30), and washer from control tube
assembly (20) protruding through bearing block (31) on
forward side of firewall (34).
1. Using care, pull control tube assembly (20) aft and
remove quadrant (29).
m. Reverse the preceding steps for reinstallation.
Rig aileron, interconnect and elevator control systems
MODEL 210 & T210 SERIES SERVICE MANUAL
6-18. ADJUSTMENT. Adjustment is accomplished
by loosening the screws, shifting tab trailing edge up
to correct for a wing-heavy condition or down to correct for a wing-light condition. Divide correction
equally on both tabs. When installing a new wing or
aileron, set tab in neutral and adjust as necessary
after flight test.
6-19.
6-1.)
I
CABLES AND PULLEYS.
(Refer to figure
6-20. REMOVAL AND INSTALLATION.
a. Remove access plates, wing root fairings and
upholstery as required.
b. Remove safety wire and relieve cable tension at
turnbuckles (5 and 8).
c. Disconnectcables from aileron bellcranks (18) and
quadrants (Index 29, figure 6-2, Sheet 2).
d. Remove cable guards and pulleys as necessary to
work cables free from aircraft.
b. Remove safety wire and relieve all cable tension at
turnbuckles (5) and (8).
c. Disconnect push-pull rods (16) at bellcranks (18).
d. (Refer to figure 6-2, Sheet 2.) Adjust turnbuckle
(33) and interconnect cable adjustment terminal (28)
nuts on interconnect cable (32) to remove slack, acquire
I
proper tension (30 pounds, ± 10 pounds), and position
both control wheels (1) 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 (5) and carry-thru cable turnbuckle (8) to
position bellcranks (18) approximately in neutral
while maintaining 40±10 pounds tension on carry-thru
cable (7).
f. Streamline ailerons with reference to flaps
(laps full UP and disregarding aileron trim tabs),
then adjust push-pull rods (16) to fit and install.
g. With ailerons streamlined, mount an inlnometer on trailing edge of aileron and set pointer to
00.
NOTE
NOTE
An inclinometer for measuring control
An inclnometer for measuring control
surface travel is available from Cessna Parts
To ease routing of cables during
reinstallation, a length of wire may be
attached to end of the cable before being
withdrawn
Leave wire
wire in
in
from aircraft.
aircraft. Leave
withdrawn from
place, routed through structure; then attach
thecable being installed, and use it to pu
cable into position. ~stop
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. Rerig aileron system in accordance with paragraph
6-17, safety turnbuckles (5) and (8), and install access
plates, fairings, and upholstery removed in step "a.".
6-17. RIGGING.
a. (Refer to figure 6-1.) Remove access plates
and upholstery as required.
Distribution (CPD 2) through Cessna Service
Stations. Refer to figure 64.
h.
Remove bar from control wheels and adjusttravel
bolts (15) to degree of travel specified in Figure 1-1.
i. Ensure all turnbuckles (5) and (8) are safetied, all
cables and cable guards are properly installed, and all
nuts are tight, and replace all parts removed for access.
WARNING
Be sure ailerons move in correct direction
when operated by the control wheels.
SHOP NOTES:
Revision 3
6-11/(6-12 blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 7
WING FLAP CONTROL SYSTEM
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
WING FLAP CONTROL SYSTEM ....
Description ...........
Operational Check ...
...
Trouble Shooting .........
Flap Motor, Transmission and
Actuator Assembly .
......
Removal and Installation .
Repair ..........
Flap Control Lever .......
Removal and Installation .
Drive Pulleys .........
..
1L3/7-1
1L3/7-1
L3/7-1
1L4/7-2
1
L5/7-3
. .1L5/7-3
1
L5/7-3
. 1L5/7-3
. . 1L5/7-3
. L5/7-3
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.
7-3. OPERATIONAL CHECK.
a. Operate flaps through their full range of travel,
Removal and Installation
Repair
..........
Bellcranks
..........
Removal and Installation
Repair ..........
Flap
s .....
.......
Removal and Insallation
Repair .
...........
Cables and Pulleys ......
Removal and Installation
Rigging .............
.
1L5/7-3
1L9/7-7
1L9/7-7
. . .1L9/7-7
.1L9/7-7
1L9/7-7
.
.1L9/7-7
1L9/7-7
. 1L9/7-7
. . 1L9/7-7
1L9/7-7
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. Check for positive shut-off of motor at the flap
travel extremes, FLAP MOTOR MUST STOP OR
DAMAGE WILL RESULT.
c. Check wing flaps for sluggish operation on the
ground with engine running.
d. 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
An inclinometer for measuring control surface
travel is available from Cessna Parts
Distribution (CPD 2) through Cessna Service
Stations. Refer to Section 6.
e. Remove access plates and attempt to rock drive
pulleys and bellcranks to check for bearing wear.
f. Inspect flap rollers and tracks for evidence of
binding and defective parts.
Revision 3
7-1
MODEL 210 & T210 SERIES SERVICE MANUAL
7-4.
TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting chart,
it may be necessary to rerig system. Refer to paragraph 7-21.
|
TROUBLE
BOTH FLAPS FAIL TO MOVE.
BINDING IN SYSTEM AS FLAPS
ARE RAISED AND LOWERED.
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 diconnected 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.
LEFT FLAP FAILS TO MOVE.
FLAPS FAIL TO RETRACT.
7-2
Revision 3
Disconnected or broken cable.
Check cable tension.
Connect or replace cable.
Disconnected push-pull rod.
Attach push-pull rod.
Disconnected or defective
UP operating switch.
Check continuity of switch.
Connect or replace switch.
Replace
MODEL 210 & T210 SERIES SERVICE MANUAL
7-4.
TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
REMEDY
FLAPS FAIL TO EXTEND.
Disconnected or defective
DOWN operating switch.
Check continuity of switch.
Connect or replace switch.
INCORRECT FLAP TRAVEL.
Incorrect rigging.
Refer to paragraph 7-21.
Defective limit switch.
Check continuity of switches.
Replace switches found defective.
7-5. FLAP MOTOR, TRANSMISSION AND ACTUATOR ASSEMBLY. (Refer to figure 7-1, sheet 2.)
7-8. FLAP CONTROL LEVER.
sheet 2.)
7-6. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position.
b. Disconnect battery cables at the battery and
insulate cable terminals as a safety precaution.
c. Remove access plates from under actuator assembly on left wing and adjacent to the drive pulleys
on both wings.
d. Relieve cable tension at turnbuckles (indexes 6,
7, 8 and 9, sheet 1.)
7-9. REMOVAL AND INSTALLATION.
a. Remove follow-up control (8) from switch mounting arm (30).
b. Remove flap operating switches (28 and 29) from
switch mounting arm (30). DO NOT disconnect electrical wiring at switches.
c. Remove knob (27) from control lever (26).
d. Remove remaining items by removing bolt (32).
Use care not to drop parts into tunnel area.
e. Reverse the preceding steps for reinstallation.
Do not overtighten bolt (32) causing lever (26) to bind.
NOTE
Remove motor (3), transmission (18), actuator assembly (17) and lower support as a unit.
e. Disconnect cables from actuator cable drive assembly (17).
f. Remove bolt (11) securing follow-up control bellcrank (10) to actuator assembly (17). Retain spacer
(9).
g. Disconnect flap motor and microswitch wiring
and tag for reference on reinstallation.
h. Remove bolts (12 and 20) securing lower support
to upper support. Retain spacer (9), bushing (19)
and washers.
i. Remove bolt (21) securing motor and transmission assembly to upper support (7).
NOTE
Although not required, nuts (2) securing motor
(3) to transmission (18) may be removed to
swing motor clear of working area for easier
removal of bolt (21).
j. Using care, work assembly out of wing through
access opening.
k. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 7-21, safety
turnbuckles and reinstall all items removed for access.
7-7. REPAIR. Repair consists of replacement of
motor, transmission or coupling. Lubricate in accordance with Section 2.
(Refer to figure 7-1,
Rig system in accordance with paragraph 7-21.
7-10.
DRIVE PULLEYS.
(Refer to figure 7-1, sheet
1.)
7-11. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position.
b. Remove access plates adjacent to drive pulley
(11).
c. Relieve cable tension at turnbuckles (7 and 8) for
removal of left hand drive pulley and relieve cable
tension at turnbuckles (6 and 9) for removal of right
hand drive pulley.
d. Remove bolt securing flap push-pull rod (17) to
drive pulley.
e. Remove bolt securing synchronizing push-pull
tube (13) to drive pulley.
f. Remove cable guards (14).
g. Remove cable lock pins (16) and disconnect
cables (10 and 18) from drive pulley. Tag cables
for reference on reinstallation.
h. Remove pivot bolt (15) attaching drive pulley to
wing structure.
i. Remove drive pulley (11) through access opening, using care not to drop bushing (12). Retain
brass washer between drive pulley and wing structure. Tape open ends of pulley to protect bearings.
j. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 7-21, safety
turnbuckles and reinstall all items removed for access.
7-3
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
1.)
BELLCRANKS.
(Refer to figure 7-1, sheet
7-14. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position.
b. Remove access plates adjacent to bellcrank (21).
I c. Remove bolt (24) securing outboard push-pull rod
(23) to bellcrank (21).
d. Remove bellcrank pivot bolt (19) and position bellcrank as necessary to expose synchronizing push-pull
tube attach point,
e. Remove bolt (22) securing synchronizing pushpull tube (13) to bellcrank (21) and work bellcrank out
through access opening using care not to drop bushing
(20). Tape open ends of bellcrank to protect needle
bearings.
NOTE
To remove synchronizing push-pull tube
(13), disconnect synchronizing push-pull
tube at bellcrank (21) and drive pulley (11).
Position synchronizing push-pull tube
through lightening holes until removal
possible through access opening.
f. Reverse the preceding steps for reinstallation. If
the outboard push-pull rod (23) and synchronizing pushpull tube (13) adjustments are not disturbed, rerigging of
the system should not be necessary. Check flap travel
and rig in accordance with paragraph 7-21, if necessary,
and
items
reinstall
removedallfor access(11)
7-15. REPAIR. Repair is limited to replacement of
bearings. Cracked, bent or excessively worn bellcranks must be replaced. Lubricate in accordance
with Section 2.
7-16.
FLAPS.
(Refer to figure 7-2.)
7-17. REMOVAL AND INSTALLATION
a. Run flaps to full DOWN position.
b. Remove access plate (7) outboard of the inboard
flap track.
c. Disconnect push-pull rod (3) at both flap attach
flap
s.
d. Remove
Remove bolt
bolt (6)
(6) at
at each
each aft
aft flap
flap track,
track, pull
pull flap
b. Relieve cable tension at turnbuckles (6, 7, 8 and
9).
c. Disconnect cables at drive pulleys (11).
d. Disconnect cables at actuator cable drive assembly (index 17, sheet 2).
e. 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.
f. Reverse preceding steps for reinstallation.
g. After cable is routed in position, install pulleys and
cable guards. Ensure cable is positioned in pulley
grooves before installing guards.
h. Rerig flap system in accordance with paragraph
7-21, safety turnbuckles, and reinstall all items removed
in step "a."
7-21 RIGGING
a. (Refer to figure 7-1, sheet 1.) Using care, run flaps
to full DOWN position.
c.
Disconnect inboard push-pull rods (17) at drive
pulleys (11).
d. Disconnected outboard push-pull rods (23) at
bellcranks (21
e. Disconnect synchronizing push-pull tubes (13)
from.If cables
are being
(1
replaced with drive pulleys
f. If cables are being replaced with drive pulleys
installed, rotate drive pulleys beyond their normal range of travel to permit cable. attachment. If
drive pulleys are not installed, it may be easier to
attach the cables prior to installing the drive pulleys
in the wings.
f. Attach the 1/8" direct cable to the forward side
of drive pulleys and the 3/32" retract cable to the aft
side of drive pulleys. (Refer to figure 7-3. )
h. Adjust synchronizing push-pull tubes (13) to
41. 87" between centers of rod end holes, tighten jamnuts and instal
Adjust inboard push-pullrods(17)to 10.81" and
outboard push-pull rods (23) to 10.39" between centers of
rod end holes, tighten jamnuts and install These
rod end holes, tightenjamnuts, and install. These
dimensions may vary in order to obtain snug fitting of
aft and remove remaining bolts. As flap is removed
flap in UP position.
from wing, all washers, rollers and bushings will fall
free. Retain these for reinstallation.
e. If the push-pull rod adjustment is not disturbed,
rerigging of the system should not be necessary. Check
flap travel and rig in accordance with paragraph 7-21, if
necessary.
j Ensure cables are properly routed and in pulley
grooves, and adjust turnbuckles to obtain specified cable
tension.
k (Refer to figure 7-1, Sheets 2 and 3.)
7-18. REPAIR. Flap repair may be accomplished
in accordance with instructions outlined in Section 18.
The ball screw assembly does not have a freewheeling feature. Therefore, the flap actuator motor MUST be shut-off at travel extremes or structural deformation will occur.
7-19.
CABLES AND PULLEYS.
(Refer to figure 7-1,
sheet 1. )
Carefully run flaps to full UP position and adjust.
7-20. REMOVAL AND INSTALLATION.
a. Remove access plates, fairings and upholstery
as required for access.
Revision 3
7-7
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 8
ELEVATOR CONTROL SYSTEM
Page No.
Aerofiche/Maual
TABLE OF CONTENTS
ELEVATOR CONTROL SYSTEM ....
2A2/8-1
Description .
........
2A2/8-1
Trouble Shooting .........
2A2/8-1
Control Column .
........
2A3/8-2
Elevators .
.......
.A3/8-2
Removal and Installation. ....
2A3/8-2
Repair
...........
2A3/8-2
8-1. ELEVATOR CONTROL SYSTEM.
figure 8- 1.)
Bellcrank ..........
2A3/8-2
Removal and Installation. . 2A7/8-6
Arm Assembly
........
2A7/8-6
Removal and Installation..
2A7/8-6
Cables and Pulleys
...
2A7/8-6
Removal and Installation . 2A7/8-6
Rigging .
.........
. 2A8/8-7
(Refer to
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.
8-2. 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
8-3.
TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting chart,
it may be necessary to rerig system. Refer to paragraph 8-14.
TROUBLE
NO RESPONSE TO CONTROL
WHEEL FORE-AND-AFT
MOVEMENT.
PROBABLE CAUSE
REMEDY
Forward or aft end of push-pull
tube disconnected.
Check visually. Attach push-pull
tube correctly.
Cables disconnected.
Check visually. Attach cables and
rig system in accordance with
paragraph 8-14.
Revision 3
s-
MODEL 210& T210 SERIES SERVICE MANUAL
8-3. TROUBLE SHOOTING (Cont).
TROUBLE
BINDING OR JUMPY MOTION
FELT IN MOVEMENT OF ELEVATOR SYSTEM.
PROBABLE CAUSE
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.
Defective control column
Check visually. Replace defective
bearing rollers.
rollers.
Defective control column
torque tube bearings.
Disconnect necessary items and
check that bearings rotate freely.
Replace defective bearing.
Control guide on aft end of ctrolsquare tube adjusted too
Loosen screw and tapered plug
in end of control tube enough to
tightly
ELEVATORS FAIL TO ATTAIN
PRESCRIBED TRAVEL.
REMEDY
l
eliminate
binding.
Defective elevator hinges.
Disconnect push-pull tube and move
elevator by hand. Replace defective hinges.
Defective pulleys or-cable
guards.
Check visally. Replace defective
parts and install guards properly.
Stops incorrectly set.
Rig in accordace 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.
e. Using care, remove elevator.
Section 6 outlines removal, installation and repair of
f. To remove left elevator use same procedure,
control column.
omitting
step "b".
g. Reverse the preceding steps for reinstallation
8-5. ELEVATORS. (Refer to figure 8-2.)
h. Set right hand elevator maintaining 0.18-inch
dimension specified in figure 8-2.
8-6.- REMOVAL AND INSTALLATION.
When
i
reinstallingbolts (13) install a washer
a. Remove stinger.
under the head of each bolt and-under each nut. Apply
b. Disconnect trim tab push-pull tube at tab actuAdhesive EA-9309 from Hysol Division, Dexter Corp.,
ator. (Refer to Section 9.)
or its equivalent, only to the shanks of bolts (13).
NOTE
Wipe off excess adhesive after installation.
8-7.
Repair
If trim system is not moved and actuator screw
8-7. REPAIR.
REPAIR
Repair may
may be
be acomplished
accomplished asas outoutis not turned, rerigging of trim system should
not be necessary after reinstallation of elevator.
I
c.
Remove bolts (13) securing torque tubes (7) to arm
assembly (8). A heat gun may be required to soften
epoxy adhesive on bolt (13).
d.
8-2
Remove bolts (6) from elevator hinges (5).
Revision 3
lined in Section 18.
inge bearings may be replaced
as necessary. IF repair has affected static balance,
check and rebalance as required.
8-8.
BELLCRANK.
(Refer to figure 8-3.)
8-9. REMOVAL AND INSTALLATION.
a. Remove access plate below bellcrank on tailcone.
MODEL 210 & T210 SERIES SERVICE MANUAL
1> 2
2
!2
212
12 ->\614.
^*^,;.\
>
H o^-
/
23
Figure 9-1.
Revision 3
n»Detail
1
20
9-4
*
-\
~21t
f^^
26
Bracket Assembly
D*'
\^
\26.
J
\," n
15.
16.
Push-PullTube
Brace
.17.Stabilizer Rear Spar
18. Support Bracket
19. Actuator
20. Sprocket
21. Chain Guard
22. Clamp
23. Chain
24. Mounting Bracket
25. Trim Tab
Sprocket (Electric Trim)
Elevator Trim Tab Control System (Sheet 2 of 3)
MODEL 210 & T210 SERIES SERVICE MANUAL
15
18
2419
21
Detail
J
BEGINNING WITH SERIAL 21062383
Figure 9-1.
Elevator Trim Tab Control System (Sheet 3 of 3)
9-5
MODEL 210 & T210 SERIES SERVICE MANUAL
* Do not overtighten nut.
1. Right Elevator
2. Trim Tab
3. Hinge Half
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
12
Spacer
Foam Filler
Horn Assembly
Bushing
Bolt
Push-Pull Tube
Hinge Pin
Screw
Nutplate
Left Elevator
13
7.
Trim
1 Sprocket-Electric
Bearing
2.
3.
4.
5.
6.
7.
9-6
5
Screw Assy.
Actuator
Sprocket Manal
Manual Trim
Trim
Sprocket Guard
Grease Zerk
Sprocket-Electric Trim
Figure 9-3.
Detail B
Elevator Trim Tab Actuator Assembly
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
10. Reverse the preceding steps for reinstallation.
11. 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).
12. Re-rig system in accordance with paragraph
9-15, safety turnbuckle (8) and reinstall all items
removed for access.
9-12. TRIM TAB FREE-PLAY INSPECTION. (Refer
to figure 9-5A.)
a. Place elevators and trim tab in neutral position
and secure from movement
b. Determine maximum allowable free-play using
the following instructions.
1. Measure chord length of extreme inboard end
of the trim tab as shown in detail A, figure 9-5A.
2. Multiply chord length by 0. 025 to obtain maximum allowable free-play.
c. Using moderate pressure, move the trim tab
trailing edge up and down by hand to check free-play.
NOTE
Measure free-play at the same point on trim
tab that chord length was measured. Total
free-play must not exceed maximum allowable. Refer to detail B, figure 9-5A.
d. If the trim tab free-play is less than the maximum allowable the system is within the prescribed
limits.
e. If the trim tab free-play is more than the maximum allowable, check the following items, for looseness while moving the trim tab up and down.
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 actuator assembly with push-pull
tube disconnected.
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 9-7 through 9-8. Recheck trim
tab free-play.
'FWD
non-faired difference between the inboard and outboard ends).
2.
Place inclinometer on trim tab, adjust inclinometer to 0* and lower tab
to degree of travel specified in figure 1-1.
3.
Position stop block (3) on cable B, maintain 0. 125 " between stop block
(3) and pulleys (4) when elevator tab is in full down position and secure
stop block (3) to cable B.
4. Raise trim tab to specified degree, place stop block (2) against stop block
(3) and secure to cable A.
5.
Place trim tab in full down position maintaining 0.125 " between stop block
(3) and pulleys (4), place stop block (1) against stop block (2) and secure to
cable B. (Recheck travel.)
Figure 9-5.
9-8
Elevator Trim Tab Travel Stop Adjustment
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
9-13. PEDESTAL COVER.
9-14. 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.
c. Remove fuel selector handle and placard.
d. Remove cowl flap handle knob.
e. Remove electric trim circuit breaker nut and
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.
9-15.
RIGGING MANUAL TRIM.
4. Tighten nut (9) and screws (13) 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
(split the non-faired difference between the inboard and outboard ends), 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 when the elevators are streamlined.
(Refer to figure
NOTE
9-1.)
CAUTION
Position a support stand under tail tiedown 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 (7 and 12).
c. Disconnect push-pull tube (15) from actuator
(19).
d. Check cable tension for 20+0-5 pounds, and
readjust turnbuckle (8) 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 (8).
e. (Refer to figure 9-4.) Rotate trim wheel (7) full
forward (nose down). Ensure position indicator (3) does
not restrict trim wheel movement. If necessary to
reposition indicator, proceed as follows:
1. Remove pedestal cover as outlined in
paragraph 9-14.
2.
Loosen nut (9) at trim, wheel pivot stud (8).
3. Loosen screws (13) securing chain guard (10)
far enough that trim wheel (7) can be moved
approximately 1/8-inch, then reposition position
indicator (3) using a thin screwdriver to pry trailing leg
of pointer out of groove in trim wheel. Reposition
position indicator as required.
9-10
Revision 3
An inclinometer for measuring control surface
travel is available from Cessna Parts
Distribution (CPD 2) through Cessna Service
Stations. Refer to Section 6.
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 (Indexes 1, 2, and 3, figure 9-5.)
as illustrated in figure 9-5 to degree of trim tab travel
specified in figure 1-1.
j. Install pedestal cover and adjust trim tab position
indicator (3) as follows:
1. Rotate trim wheel (7) to place tab at 10° up
position.
2. Locate position indicator (3) at the TAKE-OFF
triangle as viewed from the pilot seat. (Refer to step "e."
and reposition pointer if necessary.)
3.
Bend position indicator (3) as required to clear
pedestal cover. (Position indicator must NOT rub
against pedestal cover or clear cover more than 0.125inch maximum.)
k. Safety turnbuckle (Index 8, figure 9-1) 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.
MODEL 210 & T210 SERIES SERVICE MANUAL
9-16. ELECTRIC TRIM ASSIST INSTALLATION.
(Refer to figure 9-6.)
ing
9-17. DESCRIPTION. The electric elevator trim
assist installation consists of two switches mounted
on the pilot's control column, a circuit breaker
mounted on the center pedestal cover, wiring running
9-18.
aft to the electric drive assembly and a chain connect
the drive assembly to an additional sprocket
mounted on the standard manual elevator trim actuato
When the clutch (16) is not energized, the drive as
sembly "free wheels" and has no effect on manual
trim operation.
TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
SYSTEM INOPERATIVE.
Circuit breaker out.
Check visually. Reset breaker.
Defective breaker.
circuit
TRIM MOTOR OPERATING TRIM TAB FAILS TO MOVE.
REMEDY
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 spnner nuts
for proper tension.
Disconnected or broken
cable.
Operate-anual trim wheel.
Connect or replace cable.
Defective actuator.
Check actuator operation.
Replace actuator.
9.19. REMOVAL AND INSTALLATION. (Refer to
figure 9-6.)
a. Remove aft baggage compartment wall.
NOTE
Position a support stand under tail tiedown ring to prevent the tailcone from
dropping while working inside.
b. Remove cover (29) below drive assembly (6).
c. Remove cover (28) with voltage regulator
attached and carefully disconnect wiring at
connectors.
d. Remove sprocket guard (Index 5. figure 9-3)
from trim tab actuator (3).
e. Remove mounting bolts from drive assembly
and tab actuator and remove from aircraft.
f. Reverse preceding steps for reinstallation.
Check system rigging in accordance with
paragraph 9-23.
Replace
Replace
9-20. CLUTCH ADJUSTMENT. (Refer to figure 96.)
a. Remove access covers (28) & (29) below
actuator.
b. Remove safety wire and relieve cable tension
at turnbuckle (31).
c. Disconnect electric motor by unplugging the
"quick-disconnect" connectors leading to the motor
assembly.
d. Remove mounting bolts from drive assembly
(6). It is necessary to remove from stabilizer to
make the necessary adjustments to clutch.
NOTE
Step "c" isolates the motor assembly
from the remainder of the electric trim
system so it cannot be engaged during
clutch adjustment.
e.
Remove screws securing covers (17) and f 1,
3-11
MODEL 210 & T210 SERIES SERVICE MANUAL
31
REFER TO
MODEL 210 & T210 SERIES SERVICE MANUAL
32. Support Bracket
33. Screw
34. Noise Filter
31
30
34*
*NOTE
BEGINNING WITH 1980 MODEL YEAR
A-374A Noise filter must be installed
with the 400 Autopilot.
Figure 9-6. Electric Elevator Trim Assist Installation (Sheet 2 of 2)
0-13
MODEL 210 & T210 SERIES SERVICE MANUAL
1.
CTR1 Adjustment
3. Connector
electrical wiring far enough to expose the clutch
assembly.
f. Ensure the electric trim circuit breaker on the
pedestal cover is pushed in and place master
switch in the ON position.
g. Operate control wheel-mounted trim switch (3)
UP or DOWN to energize the solenoid clutch (16).
h. Attach the spring scale to chain and pull scale
slowly until slippage is noted.
Repeat Steps "g" and "h" several times to
i.
break the initial friction of the clutch.
Repeat Step "h" very slowly, carefully
j.
watching the indicator on the spring scale.
Slippage should occur between 38.6 to 42.5 lbs.
k. IF tension is not within tolerance, loosen
OUTSIDE spanner nut (14) which acts as a lock.
Tighten INSIDE spanner nut to increase clutch
tension and loosen nut to decrease clutch tension.
When clutch slippage torque is within
l.
tolerance (step j"). then tighten outside spanner
nut against inside nut.
m. Connect electrical wiring to motor assembly
which was removed in Step "c" re-rig
in accordance with paragraphs 9-15 and 9-24 and
reinstall all items removed for access.
RED and BLACK wire leading to the motor assembly.
CAUTION
Ensure CTR adjustments (Index 1 and 2,
Figure 9-7) are both turned fully CCW
to limit initial voltage to motor and
voltmeter.
d. Using 18 ga. jumper wires or equivalent, 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 CTR2 adjustment screws on
the voltage regulator counterclockwise (CCW), then
slowly turn adjustment screws clockwise (CW) until
a 11 volt output is obtained for both (RED and BLACK)
lead.
to see if full "NOSE UP" to full "NOSE
Check
system
g. trim
" to full "NOSE UP"
DOWN
DOWN" and full "NOSE
is 39± 1 seconds.
h. Remove voltmeter and reconnect the motor
9-21.
VOLTAGE REGULATOR ADJUSTMENT.
power leads. Be sure to connect RED to
assemblyis 39±1 seconds.
(Refer
to figure
9-6.)
all items
removed for access.
reinstall
RED and BLACK to BLACK when reconnecting leads.
a. Remove access cover (27)
i. Check trim system for proper operation and
b. Connect an external power source of 27. 5 volts
all items removed for access.
reinstall
if
or
system,
electrical
aircraft
the
to
dc continuous
an external power supply is not available, run the
aircraft engine at approximately 1000 RPM to maintain the normal operating aircraft voltage.
c. Disconnect the electrical power leads to the
9-14
CAUTION
The trim motor should be allowed to cool
MODEL 210 & T210 SERIES SERVICE MANUAL
TRIMTAB
ANGLE
ANGLE
HORN ASSEMBLY
CABLE
*
1
WEIGHT
(14 to 22 lbs total)
Figure 9-8.
Trim Tab Simulated Air Load Test
between voltage regulator adjustments
approximately 5 minutes if several actuations of the motor becomes necessary during adjustment.
9-22. TRIM TAB SIMULATED AIR LOAD TEST.
(Refer to figure 9-8.)
NOTE
The manual elevator trim control system
must be properly rigged, the aircraft
electrical operating voltage must be normal, the electric trim assist clutch must
be properly adjusted and the elevator must
be in neutral position prior to completing
the following steps.
a. Attach two angles approximately 18 inches in
length to the trailing edge of the trim tab with clamps
as illustrated to prevent bending of tab trailing edge.
b. Attach a cable directly aft of the trim tab horn
assembly.
c. Attach 14 pounds minimum to 22 pounds maximum of weight (including the angles, clamps and
cable) to the cable and operate the trim switch to
place the tab in the UP position. The clutch MUST
lift 15 pounds weight to the FULL UP position but
must slip at 18 pounds.
NOTE
If the electric trim clutch slips prior to
lifting the required weight to the full up
position, DO NOT READJUST CLUTCH,
refer to step "d" or step 5 to locate and
remove thereason for excessive friction
in the elevator trim control system.
d. Check the trimtab hinge and linkage for binding,
check the trim system cables and chains for proper
tension, check system pulleys and actuator for binding.
e. After the trim system has been thoroughly
checked and excessive friction removed, repeat
step "c", or step 3.
9-l5
MODEL 210 & T210 SERIES SERVICE MANUAL
9-23. RIGGING - ELECTRIC TRIM ASSIST. (Refer
to figure 9-6.)
a. The standard manual elevator trim control
system MUST be rigged in accordance with paragraph
9-15 prior to rigging the electric trim assist.
b. Move elevator trim tab to full "NOSE UP"
position.
SHOP NOTES:
9-16
c. Remove access cover (29) located in under
side of right stabilizer.
d. Locate turnbuckle (31) terminal point 0.75
inches from drive assembly housing and adjust until
chain deflection between sprockets is approximately
0.25 inches.
e. Resafety turnbuckle and reinstall all items
removed for access.
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 10
RUDDER CONTROL SYSTEM
Page No.
Aerofiche/Manual
TABLE OF CONTENTS
RUDDER CONTROL SYSTEM .....
......
Description
.......
Trouble Shooting .
Rudder Pedal Assembly .....
Removal and Installation .
Rdder ............
2B13/10-1
2B13/10-1
.2B13/10-1
.2B17/10-5
. 2B17/10-5
2B17/10-5
.. 2B17/10-5
Removal and Installation
2B17'10-5
.....
Repair
2B17/10-5
.....
Cables and Pulleys .
Removal and Installation . . 2B17/10-5
. 2B20/10-8
......
.
.
Rigging
10-1. RUDDER CONTROL SYSTEM.
ure 10-1.)
(Refer to fig-
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 through 1977 models.
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.
TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting chart,
it may be necessary to rerig 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
cables.
Revision 3
10-1
MODEL 210 & T210 SERIES SERVICE MANUAL
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. Replace defective
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 10-1 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
Repair
MODEL 210 & T210 SERIES SERVICE MANUAL
25
NOTE
23
Brake links (5), bellcranks (22), brake torque
tubes (19) and attaching parts for the RIGHTHAND rudder pedals are replaced with hubs (8)
when dual controls are NOT installed. These
hubs are attached to each end of the forward
rudder
bars.
25*
2
24
* THRU 21064806
4
BEGINNING WITH 21064807
CLEARANCE
1. Anti-Rattle Spring
2. Pedal
3. Shaft
4. Spacer
5. Brake Link
6. Cable (Left Forward)
7. Cable (Right Forward)
8. Single Controls Hub
9. Pin (Stowable Pedals Only)
10. Stowable Pedals Controls
11. Bearing Block
12. Right Rudder Cable Arm
13. Left Rudder Cable Arm
14. Aft Rudder Bar
15. Nosewheel Steering Arm
16. Rudder Trim Bungee Arm
17. Forward Rudder Bar
18. Master Cylinder
19. Brake Torque Tube
20. Bracket
21. Bearing
22. Bellcrank
23. Washer
24. Pedal Extension
25.
Shaft
Detail A
13
17
15
CLEARANCE
HOLE FORWARD
18
23
20
Detail
C
At least one washer (23) must be
installed at the location shown.
Figure 10-2.
10-4
19
NOTE
Detail B
STOWABLE RUDDER
PEDALS INSTALLATION
THRU 1977 MODELS
21
Rudder Pedal Installation
MODEL 210 & T210 SERIES SERVICE MANUAL
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Steering Arm
Steering Bungee
Adjustable Rod End
Whiffletree (Steering Bellcrank)
Link Rod Assembly
Clamp
10
Boot
Boot Retainer
Right Rudder Bar Arm
Left Rudder Bar Arm
7
4
Figure 10-3.
10-4.
Nose Gear Steering Installation
RUDDER PEDAL ASSEMBLY.
10-5. REMOVAL AND INSTALLATION. (Refer to
figure 10-2.)
a. Remove carpeting, shields and soundproofing
from the rudder pedal and tunnel areas as necessary
for access.
b. Disconnect brake master cylinders (18) and
parking brake cables at pilot's rudder pedals.
c. Remove rudder pedals (2) and brake links (5).
d. Disconnect stowable rudder pedal controls (10).
e. Remove fairing from either side of vertical fin,
remove safety wire and relieve cable tension by loosening turnbuckles (index 10, figure 10-1).
f. Disconnect cables (6 and 7) from rudder bar
arms (12 and 13).
g. Disconnect rudder trim bungee from rudder bar
arm (16).
h. (Refer to figure 10-3.) Disconnect whiffletree
link rod assemblies (5) at rudder bar arms (9 and 10).
i. (Refer to figure 10-2.) Remove bolts securing
bearing blocks (11) and carefully work rudder bars
out of tunnel area.
NOTE
The two inboard bearing blocks contain clearance holes for the rudder bars at one end and
a bearing hole at the other. Tag these bearing blocks for reference on reinstallation.
j. Reverse the preceding steps for reinstallation.
Lubricate rudder bar assemblies as outlined in Section 2. Rig system in accordance with paragraph
10-11, safety turnbuckles and reinstall all items removed for access.
10-6.
RUDDER.
(Refer to figure 10-4.)
10-7. REMOVAL AND INSTALLATION.
a. Remove stinger.
b. Disconnect tail navigation light wire.
c. Remove fairing from either side of vertical fin,
remove turnbuckles (index 10, figure 10-1.)
d. Disconnect cables (4 and 6) from rudder bellcrank (3).
e. With rudder supported, remove all hinge bolts
(2) and using care, lift rudder free of vertical fin.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 10-11,
safety turnbuckles and reinstall all items removed
for access.
10-8. REPAIR. Repair may be accomplished as
outlined in Section 18.
10-9. CABLES AND PULLEYS.
10-1.)
(Refer to figure
10-10. REMOVAL AND INSTALLATION
a. Remove seats, upholstery and access plates as
necessary.
10-,
MODEL 210 & T210 SERIES SERVICE MANUAL
(2 x 4)
VERTICAL FIN
RUDDER
BLOCK
BLOCK RUDDER
HALF THE
DISTANCE BETWEEN
STRAIGHTEDGES
WIRE POINTER
MEASURING
RUDDER
TRAVEL
ESTABLISHING NEUTRAL
POSITION OF RUDDER
1.
Establish neutral position of rudder by clamping straightedge (such as wooden 2 x 4) on each side of
fin and rudder and blocking trailing edge of rudder half the distance between straightedges as shown.
2.
Tape a length of soft wire to the stinger in such a manner that it can be bent-to index at the lower
corner of the rudder trailing edge.
3.
Using soft lead pencil, mark rudder at point corresponding to soft wire indexing point (neutral).
4.
Remove straightedges and blocks.
5.
Hold rudder against right, then left, rudder stop. Measure distance from pointer to pencil mark
on rudder in each direction of travel. Distance should be between 8.12" and 8.72".
Figure 10-5.
Checking Rudder Travel
10-7
MODEL 210 & T210 SERIES SERVICE MANUAL
b. Remove safety wire, relieve cable tension and
disconnect cables at turnbuckles (10).
c. Disconnect cables (3 and 4) at rudder bar arms.
d. Remove cable guards, pulleys and fairleads as
necessary to work cables free of aircraft.
NOTE
To ease routing of cables, a length of wire
may be attached to end of the cable before
being withdrawn from aircraft. Leave
wire in place, routed through structure;
then attach cable being installed and pull
the cable into position,
e. Reverse the preceding steps for reinstallation.
f. After cable is routed in position, install pulleys,
fairleads and cable guards. Ensure cable is positioned in pulley grooves before installing guards.
g. Re-rig system in accordance with paragraph 1011, safety turnbuckles and reinstall all items removed in step "a".
10-11. RIGGING.
a. Remove fairing from either side of vertical fin,
remove safety wire and relieve cable tension at turnbuckles (index 10, figure 10-1).
b. Open landing gear doors. (Refer to Section 5. )
c. Tie down or weight tail to raise nosewheel free
of ground.
d.
Extend strut and ensure nose gear is centered
against the external centering lug. (Neutral position. )
e. (Refer to figure 10-3.) Disconnect steering bungee adjustable rod end (3) from whiffletree (4).
SHOP NOTES:
10-8
f. Remove pedestal cover in accordance with Section 9.
g. Remove lower pedestal panel (index 14, figure
9-4).
h. Disconnect rudder trim bungee from rudder bar
arm (index 16, figure 10-2).
i. Clamp rudder pedals in neutral position.
j. Adjust turnbuckles (index 10, figure 10-1) to
streamline rudder with 30±10 lbs tension on cables.
k. Remove clamps from rudder pedals.
1. Adjust travel stop bolts (index 13, figure 10-1)
to obtain degree of travel specified in figure 1-1.
Figure 10-5 illustrates correct travel and one method
of checking.
m. Adjust length of rod end (3) to align with whiffletree (4) and install bolt. DO NOT PRELOAD BUNGEE.
n. Connect rudder trim bungee and rig trim system
as outlined in Section 11.
o. Operate rudder system, checking for ease of
movement and full travel. Check cable tension with
rudder in various positions. Cable tension should
not be less than 20 pounds or more than 40 pounds
in any position.
p. Check that all turnbuckles are safetied and reinstall all items removed for access.
q. Lower nosewheel to ground.
_WARNING
WARNN
Be sure rudder moves in the correct direction when operated by the rudder pedals.
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 11
RUDDER TRIM CONTROL SYSTEM
Page No.
Aerofiche/Manual
TABLE OF CONTENTS
RUDDER TRIM CONTROL SYSTEM
Description ..................
TroubleShooting ............
Removal and Installation of
System Components ........
Indicator Assembly ........
2C1/11-1
2C1/11-1
2C1/11-1
Wheel and Gear Box
Assembly ...............
ChainAssembly ...........
Gimbal Assembly .........
Bungee Assembly .........
Rigging Rudder Trim System ..
2C3/11-3
2C3/11-3
11-1. RUDDER TRIM CONTROL SYSTEM.
figure 11-1.)
(Refer to
11-2. DESCRIPTION. The rudder trim system is
comprised of a trim control wheel and gear box
assembly located in the upper control pedestal, which
is connected by a chain assembly to a gimbal assembly in the lower pedestal. The gimbal assembly is
11-3.
2C3/11-3
2C3/11-3
2C3/11-3
2C3/11-3
2C3/11-3
attached to a stop bracket, which is attached to the
rudder trim bungee. The bungee's push-rod assembly is attached to the right-hand rudder bar assembly. The rudder control system, rudder trim control
system, and the nosewheel steering system are interconnected and adjustments to any one system will
affect the others.
TROUBLE SHOOTING.
NOTES
This trouble shooting chart should be used in conjunction with
the chart shown in Section 10.
Due to remedy procedures in the following trouble shooting chart,
it may be necessary to rerig system. Refer to paragraph 11-5.
TROUBLE
FALSE READING ON TRIM
POSITION INDICATOR.
HARD OR SLUGGISH OPERATION OF TRIM WHEEL.
FULL TRIM TRAVEL
NOT OBTAINED.
PROBABLE CAUSE
REMEDY
Improper rigging.
Refer to note above.
Worn, bent or disconnected
linkage,
Check visually. Repair or
replace parts as necessary.
Worn, bent or binding linkage.
Check visually. Repair or
replace parts as necessary.
Incorrect rudder cable tension.
Check and adjust rudder cable
tension.
Rudder trim system improperly
rigged.
Refer to note above.
Revision 3
1-1
MODEL 210 & T210 SERIES SERVICE MANUAL
1.
2.
3.
4.
5.
6.
7.
8.
* THRU 1981 MODELS
* BEGINNING WITH 1982 MODELS
Chain Guard
Pedestal Assembly
Upper Panel
Lower Panel
Bearing Bracket
Gimbal Half Assembly
Bearing Bracket
Sprocket Drive Nut
5*
7*
9. Shim
10. Gimbal Cover Plate
A
11. Stop Bracket
12. Left-Hand Chain Guard
13. Bungee
19
21
14. Idler Sprocket
16. Washers
17. Support Assembly
18. Gear Box Assembly
19. Mounting Bracket
20. Bushing
rm Assembly
21. Indicator a
22. Trim Wheel
23. Washers
24. Dual Sprocket Assembly
25. Spacer
26. Sprocket Support
27. Chain
28. Right-Hand Chain Guard
28
Detail
A
NOTE
Lubricate bungee screw and
sprocket drive nut threads
accordance with Section 2.
Figure 11-1. Rudder Trim Control System
11-2
11-2
6*
7*
5*
MODEL 210 & T210 SERIES SERVICE MANUAL
11-4. REMOVAL AND INSTALLATION OF SYSTEM
COMPONENTS. (Refer to figure 11-1.)
a. INDICATOR ASSEMBLY.
1. Remove pedestal cover in accordance with
procedures outlined in Section 9.
2. Remove four screws attachingmounting
bracket assembly (19) to pedestal assembly (2).
3. Remove indicator assembly as a unit.
4. Reverse preceding steps for installation.
h. WHEEL AND GEAR BOX ASSEMBLY.
1. Remove pedestal cover as outlined in Section
9.
Loosen chain (27) by loosening belt securng
2
idler sprocket (14) and sliding sprocket inboard in
slot in supportangle (15).
3. Remove upper panel (3) and disconnect chain
(27) at connecting link.
4. Remove four bolts attaching gear box assembly (18) to pedestal assembly (2).
5. Remove bolts attaching idler sprocket (14)
and chain guards (12) and (28).
6. Remove wheel and gear box assembly a a
unit
NOTE
If wheel and gear box assembly is disassembled, install washers (16) and (23) as
required to nest sprockets and prevent end
play.
7. Reverse preceding steps for installation.
c. CHAIN ASSEMBLY.
1. Remove pedestal cover as outlined in Section
9.
2. Remove upper panel (3).
3. Remove access cover directly below-and aft
of pedestal in floor.
4. Remove fuel selector shaft, then remove
lower panel (4).
5. Loosen chain (27) by loosening bolt securing
idler sprocket (14) and sliding sprocket inboard in
slot in support angle (15).
6. Disconnect chain at connecting link.
7. Remove bolt attaching bungee (13) to stop
bracket (11).
8.
Pull gimbal assembly (items 5, 6, 7, 8, 9,
1. and 11) aft awav from bungee (13).
9. Remove chain (27) from sprocket drive nut
10. Reverse preceding steps for installation.
d. GIMBAL ASSEMBLY.
1. Remove pedestal cover as outlined in Section
.
2. Remove access cover directly below and aft
of pedestal in floor.
3. Remove fuel selector shaft, then remove
lower panel (4).
4. Loosen chain (27) by loosening bolt securing
idler sprocket (14) and sliding sprocket inboard in
slot in support angle (15).
5. Disconnect chain at connecting link.
6. Remove bolt attaching bungee (13) to stop
bracket (11).
7. Pull gimbal assembly (items 5, 6, 7, 8, 9,
10 and 11) aft; remove from aircraft.
NOTE
If gimbal assembly is to be diassembled,
upon reassembly, shims (9) should be nstalled between gimbal half assembly (6)
and cover plate assembly (10) to maintain
.002 to. 04-inch end play on sprocket.
8. Reverse preceding steps for installation.
e. BUNGEE ASSEMBLY.
1. Remove pedestal cover as otlined in Section
9.
2. Remove upper panel(3).
3. Remove access cover directly below and aft
of pedestal in floor.
4. Remove fuel selector shaft, then remove
lower panel (4).
Loosen chain (27) by loosening bolt securing
5.
idler sprocket (14) and sliding sprocket inboard in
slot in support angle (15).
Disconnect chain at connecting link.
6
7. Remove bolts attaching idler sprocket (14)
and chain guards (12) and (28) to support angle (15).
8. Remove bolts attaching chain guard to stop
bracket (11); remove chain guards.
9. Remove bolt attaching bungee (13) to stop
bracket (11).
10. Pull gimbal assembly (items 5, 6, 7, 8, 9,
10 and 11) aft; remove from aircraft.
11. Disconnect bungee push-rod assembly from
right-hand rudder bar assembly.
12. Using care, remove bungee from tunnel
area, aft, through pedestal.
13. Reverse preceding steps for installation.
NOTE
Upon installation, lubricate bungee screw
and sprocket drive nut threads per Section 2.
11-5. RIGGING RUDDER TRIM SYSTEM. (Refer to
Figure 11-1.)
NOTE
Rudder control system and nose wheel
steering system must be correctly rigged
prior to rigging the rudder trim system.
b. Remove upper pedestal panel
c. Remove access cover directly below and aft of
pedesl in floor.
d. Remove fuel selector shaft, then remove lower
pedestal panel
e. Loosen chain by loosening bolt securing idler
sprocket, and sliding sprocket inboard in slot in support angle; disconnect chain.
f. Remove bolt attaching bungee to stop bracket;
unscrew gimbal assembly from actuator drive screw.
g. Disconnect bungee push-pull rod from right-hand
rudder bar assembly.
h. Tie down or weight tail to raise nose wheel free
of ground.
i. Ensure rudder pedals and rudder are in neutral
position.
11-3
MODEL 210 & T210 SERIES SERVICE MANUAL
. Attach bungee push-pull rod to right-band rudder
bar assembly.
. Install lower panel assembly and bearing brackets.
1. Screw gimbal assembly onto bungee drive screw
until studs on gimbal half assembly align with holes
in bearing brackets and nutplate on stop bracket aligns
with approximate center of slot in bungee stop arm.
m. Install and tighten bolts, washers and nuts.
n. String chain over idler sprocket and sprocket in
wheel and gear box assembly; connect chain at con-
necting link.
NOTE
Indicator assembly should be installed with
SHOP NOTES:
11-4
Revision
rudder pedals in neutral position. If indicator does not line up with centerline of aircraft, bend indicator left or right as required.
o. Tighten chain by moving idler sprocket outboard
in slot in support angle.
p. Install full selector shaft
q. Install upper paneL
r. Install floor access covers and pedestal cover.
s. Remove blocking from rudder and pedals.
t. Lower aircraft
WARNING
Be sure rudder moves in correct direction
when operated by the trim control wheel.
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 12
ENGINE
(NORMALLY ASPIRATED)
REFER TO SECTION 12A FOR TURBOCHARGED
WARNING
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the propeller,
since a loose or broken wire or a component
malfunction could cause the propeller to rotate.
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
ENGINE COWLING ..........
2C15/12-2A
Description ..........
.2C15/12-2A
Removal and Installation .....
2C15/12-2A
Cleaning and Inspection .....
. 2C15/12-2A
Repair .
.............
2C15/12-2A
Cowl Flaps
...........
2C15/12-2A
Description ........
.2C15/12-2A
Removal and Installation . . .2C15/12-2A
Rigging ..........
.2C15/12-2A
ENGINE ..............
.2C15/12-2A
Description ..........
2C15/12-2A
Engine Data .........
. 2C16/12-3
Time Between Overhaul (TBO) . . . 2C17/12-4
Overspeed Limitations .....
.2C17/12-4
Trouble Shooting .
.....
..2C18/12-5
Static Run-Up Procedures . . . . . 2C20/12-7
Removal ...........
.2C20/12-7
Cleaning
.........
Accessories Removal. .....
Inspection ..........
Buildup
.........
Installation ..........
Flexible Fluid Hoses ..
..
Pressure Test . ....
Replacement ........
..
. .2C22/12-9
.2C22/12-9
.. 2C22/12-9
..2C22/12-9
.2C22/12-9
. . 2C24/12-11
. .2C24/12-11
.2C24/12-11
Engine Baffles .........
. 2D1/12-12
Description .....
2D1/12-12
Cleaning and Inspection . . . .2D1/12-12
Removal and Installation . . . 2D1/12-12
Repair
..
.....
ENGINE OIL SYSTEM .......
Description ........
Trouble Shooting .......
.2D1/12-12
.2D1/12-12
2D1/12-12
. 2D2/12-13
Full-FlowOilFilter ...........
Description ...............
Removal and Installation
(FilterElement) ..........
Full-Flow Oil Filter
(Beginning with Serial
21064136) ...............
Description .............
Removal ...............
Installation .............
Filter Adapter. 210 Thru Serial
21064780; T210 Thru Serial
21064781 ...................
Removal ..................
2D4/12-16
2D4/12-16
2D4/12-16
2D6/12-18
2D6/12-18
2D6/12-18
2D7/12-18A
2D7/12-18A
2D7/12-18A
Disassembly, Inspection,
and Reassembly ..........
Installation ...............
Filter Adapter. 210, Beginning
with 21064781; T210,
Beginning with 21064782 ....
Oil Cooler ...................
Description ...............
ENGINE FUELSYSTEM ............
Description ..................
Fuel-Air Control Unit .........
Description ...............
Removal and Installation ...
2D7/12-18A
2D9/12-20
2D9/12-20
2D9/12-20
2D9/12-20
2D9/12-20
2D9/12-20
2D10/12-21
2D10/12-21
2D11/12-22
Cleaning and Inspection ....
2D11/12-22
Adjustments
2D11/12-22
..............
Fuel Manifold Valve ..........
Description ...............
2D11/12-22
2D11/12-22
Removal ..................
Cleaning ..................
Installation ...............
Fuel Discharge Nozzles .......
2Dl/12-22
2D11/12.22
2D12/12-23
2D12/12-23
Revision 3
12-1
MODEL 210 & T210 SERIES SERVICE MANUAL
2D12/12-23
..........
Removal .
. 2D12/12-23
Cleaning and Inspection .
2D12/12-23
Installation .........
.. 2D12/12-23
.
Fuel Injection Pump ..
2D12/12-23
Description .........
2D13/12-24
Removal ..........
2D13/12-24
Installation .........
2D13/12-24
Adjustment .......
Auxiliary Electric Fuel Pump Flow
.... 2D14/12-25
Rate Adjustment ....
.... 2D14/12-25
INDUCTION AIR SYSTEM ...
2D14/12-25
Description ...........
2D14/12-25
Airbox .............
. . . 2D14/12-25
Removal and Installation
. 2D14/12-25
Cleaning and Inspection .
.2D14/12-25
Induction Air Filter. .......
2D14/12-25
Description .........
. . . 2D14/12-25
Removal and Installation
2D14/12-25
Cleaning and Inspection ....
2D14/12-25
IGNITION SYSTEM ..........
2D14/12-25
Description ..........
... 2D16/12-27
Trouble Shooting ......
2D17/12-29
Magnetos ............
2D17/12-29
Description .........
..
. .. 2D17/12-29
Removal .....
2D17/12-29
Internal Timing .......
Installation and Timing-toEngine ..........
2D17/12-29
2D18/12-30
Maintenance ........
SHOP NOTES:
12-2
.2D19/12-31
Magneto Check .......
2D20/12-32
Spark Plugs ...........
2D20/12-32
ENGINE CONTROLS .........
2D20/12-32
Description ..........
2D20/12-32
Rigging .............
2D20/21-32
Throttle Control .......
2D21/12-33
Mixture Control .......
Throttle-Operated Microswitch. 2D21/12-33
Landing Gear Warning Horn . . 2D21/12-33
...... 2D23/12-35
Propeller Control
2D23/12-35
STARTING SYSTEM .........
2D23/12-35
Description ...........
2D23/12-35
Trouble Shooting .........
2D24/12-36
Primary Maintenance .......
2D24/12-36
..
........
Starter Motor
Removal and Installation . . . 2D24/12-36
2D24/12-36
EXHAUST SYSTEM ..........
2D24/12-36
Description ...........
Economy Mixture Indicator (EGT). . 2D24/12-36
2D24/12-36
Removal and Installation .....
2D24/12-36
Inspection
.......
EXTREME WEATHER MAINTENANCE . 2E2/21-38
2E2/12-38
Cold Weather ..........
2E2/21-38
Hot Weather ...........
. .2E2/21-38
Seacoast and Humid Areas. ..
2E2/12-38
Dusty Areas ...........
2E2/12-38
Ground Service Receptacle ...
MODEL 210 & T210 SERIES SERVICE MANUAL
12-1.
ENGINE COWLING.
12-2. DESCRIPTION. The engine cowling is divided
into four major removable segments. The left upper
cowling segment has two access doors, one at the upper front provides access to the oil filler neck and
one at the left aft side provides access to the oil dipstick. The right and left nose caps are fastened to
the lower engine nacelle and to each other with
screws. The right and left upper cowl segments are
secured with quick-release fasteners and either segment may be removed individually. The lower engine nacelle is an extension of the fuselage and provides fairing for the nose wheel in its retracted posttion.
12-3. REMOVAL AND INSTALLATION.
a. Release the quick-release fasteners attaching
the cowling to the fuselage and at the parting surfaces
of the left and right segments.
b. Remove screws securing the left and right nose
cap together and to the lower engine nacelle.
c. Disconnect air ducts from nose caps and remove
caps.
d. Reverse the preceding steps for reinstallation.
Ensure the baffle seals are turned in the correct
direction to confine and direct air flow around the
engine. The vertically installed seals must fold
forward and the side seals must fold upwards.
12-4. CLEANING AND INSPECTION. Wipe the inner surfaces of the cowling segments with a clean
cloth saturated with cleaning solvent (Stoddard or
equivalent). If the inside surface of the cowling is
coated heavily with oil or dirt, allow solvent to soak
until foreign material can be removed. Wash painted
surfaces of the cowling with a solution of mild soap
and water and rinse thoroughly. After washing, a
coat of wax may be applied to the painted surfaces to
prolong paint life. After cleaning, inspect cowling
for dents, cracks, loose rivets and spot welds. Repair all defects to prevent spread of damage.
12-5. REPAIR. If cowling skins are extensively
damaged, new complete sections of the cowling
should be installed. Standard insert-type patches
may be used for repair if repair parts are formed
to fit contour of cowling. Small cracks may be stopdrilled and small dents straightened if they are reinforced on the inner surface with a doubler of the
same material as the cowling skin. Damaged reinforcement angles should be replaced with new parts.
Due to their small size, new reinforcement angles
are easier to install than to repair the damaged part.
12-6.
COWL FLAPS.
12-7. DESCRIPTION. Cowl flaps are provided to
aid in controlling engine temperature. Two cowl
flaps, operated by a single control in the cabin, are
located at the lower aft end of the engine nacelle.
The engine exhaust tailpipes extend through cutouts
in the aft portion of each cowl flap.
12-8. REMOVAL AND INSTALLATION.
(See
figure 12-1.)
a. Place control lever (2) in the OPEN position.
b. Disconnect control cevises (12) from shockmounts (13).
c. Remove safety wire securing hinge pins (9) to
cowl flaps, pull pins from hinges and remove flaps.
d. Reverse the preceding steps for reinstallation.
Rig cowl flaps, if necessary, in accordance with
paragraph 12-9.
12-9. RIGGING. (See figure 12-1.)
a. Disconnect control clevises (12) from shockmounts (13).
b. Check to make sure that the flexible controls
reach their internal stops in each direction. Mark
controls so that full travel can be readily checked
and maintained during the remaining rigging procedures.
c. Place control lever (2) in the CLOSED position.
If the control lever cannot be placed in the closed
position, loosen clamp (5) at upper end of controls
and slip housings in clamp or adjust controls at
upper clevis (4) to position control lever in bottom
hole of position bracket (3).
d. With the control lever in CLOSED position, hold
one cowl flap closed (against the rubber bumpers on
the fuselage), loosen jam nut and adjust clevis (12) on
the control to hold cowl flap in this position and install bolt.
NOTE
If the lower control clevis (12) cannot be adjusted far enough to streamline flap and still
maintain sufficient thread engagement, loosen
the lower control housing clamp (8) and slide
housing in clamp as necessary. Be sure
threads are visible in clevis inspection holes.
e. Repeat the preceding step for the opposite cowl
flap. Cowl flaps should open approximately 5.00
inches when measured in a straight line from the aft
edge of door to firewall.
g. Check that all clamps and jam nuts are tight.
12-10.
ENGINE.
12-11. DESCRIPTION. An air cooled, wet-sump,
six-cylinder, horizontally-opposed, direct-drive,
fuel injected, Continental IO-520-L series engine
driving a constant-speed-propeller is used to power
the aircraft. The cylinders, numbered from rear to
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 numbers 2, 4 and 6.
Refer to pargraph 12-12 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 Supply Division.
Revision 2
12-2A/12-28 Blank
MODEL 210 & T210 SERIES SERVICE MANUAL
12-12.
ENGINE DATA.
Aircraft Series
210
Model (Continental)
IO-520-L
BHP Maximum for Take-Off
(5 Minutes) at RPM
BHP Maximum Except Take-Off
RPM (Max. Continuous)
300
2850
285
2700
Number of Cylinders
6-Horizontally Opposed
Displacement
Bore
Stroke
520 Cubic Inches
5.25 Inches
4.00 Inches
Compression Ratio
Magnetos
8.5:1
Slick Model 662 thru 1979 Models
Slick Model 6210 Begining with 1980 Models
Fires 22 ° BTC Upper Right
and Lower Left
Fires 22 ° BTC Upper Left
and Lower Right
Right Magneto
Left Magneto
Firing Order
Spark Plugs
Torque
Fuel Metering System
Unmetered Fuel Pressure
Nozzle Pressure
1-6-3-2-5-4
18mm (Refer to Continental Service Bulletin
M77-10 for factory approved spark plugs
and required gap)
330 30 LB-IN.
Continental Fuel Injection
9.0 to 11.0 PSI at 600 RPM
31.0 to 33.0 PSI at 2850 RPM
3.5 to 4.0 PSI at 600 RPM
17.5 to 18.5 PSI at 2850 RPM
Oil Sump Capacity
With External Filter
10 U.S. Quarts
11 U.S. Quarts
Tachometer
Mechanical Drive
Oil Pressure (PSI)
Minimum Idling
Normal
Maximum (Cold Oil Starting)
Connection Location
10
30 to 60
100
Between No. 2 and No. 4 Cylinders
Oil Temperature
Normal Operating
Maximum Permissible
Probe Location
Within Green Arc
Red Line (240°F)
Below Oil Cooler
Cylinder Head Temperature
Normal Operating
Maximum
Probe Location
Economy Mixture Indicator (EGT)
Probe Location
Approximate Dry Weight
Within Green Arc
Red Line (460'F)
Lower Side of Number 3 Cylinder
Lower Side of Number 1 Cylinder
Lower Side of Number 4 Cylinder
Without A/C
Lower Side of Number 1 Cylinder
With A/C
Exhaust Collector L.H. Side
thru 21062273
21062274 &on
21064064 & on
21064064 & un
471 LB. (Weight is approximate and will vary with
optional accessories installed.)
12-3
MODEL 210 & T210 SERIES SERVICE MANUAL
12-12A. TIME BETWEEN OVERHAUL (TBO). Teledyne Continental Motors recommends engine overhaul
at 1700 hours operating time for the IO-520-L series
engines. Refer to Continental Aircraft Engine Service
Bulletin M81-22, and to any superseding bulletins,
revisions or supplements thereto, for further recommendations. At the time of overhaul, engine accessories should be overhauled. Refer to Section 14 for
propeller and governor overhaul periods.
12-12B. OVERSPEED LIMITATIONS. The engine
must not be operated above specified maximum continuous RPM. However, should inadvertant overspeed occur, refer to Continental Aircraft Engine
Service Bulletin M75-16, and to any superseding
bulletins, revisions or supplements thereto, for
further recommendations.
Detail A
Detail A
THRU 21064535
BEGINNING WITH 21064536
MODEL 210 & T210 SERIES SERVICE MANUAL
12-13.
TROUBLE SHOOTING.
TROUBLE
ENGINE FAILS TO START.
PROBABLE CAUSE
REMEDY
Improper use of starting
procedure.
Refer to Pilot's Operating Handbook
Defective aircraft fuel system.
Refer to Section 13.
Spark plugs fouled.
Remove and clean. Check gaps and
insulators. Use new gaskets. Check
cables to persistently fouled plugs.
Defective magneto switch or
grounded magneto leads.
Check continuity, repair or replace
switch or leads.
Defective ignition system.
Refer to paragraph 12-79.
Excessive induction air leaks.
Check visually.
air leaks.
Dirty screen in fuel control unit
or defective fuel control unit.
Check screen visually. Check fuel
flow through control unit. Replace
defective fuel control unit.
Defective electric fuel pump.
Refer to Section 13.
Defective fuel manifold valve
or dirty screen.
Check fuel flow through valve.
Remove and clean. Replace if
defective.
Clogged fuel injection lines or
discharge nozzles.
Check fuel through lines and nozzles.
Clean lines and nozzles. Replace if
defective.
Fuel pump not permitting fuel
from auxiliary pump to bypass.
Check fuel flow through engine-driven
fuel pump. Replace engine-driven
pump.
Vaporized fuel in system.
Refer to Pilot's Operating Handbook
Fuel tanks empty.
Visually inspect tanks. Fill with
proper grade and quantity of gasoline.
Fuel contamination or water in
fuel system.
Open fuel strainer drain and check
for water. Drain all fuel and flush
out fuel system. Clean all screens,
fuel lines, strainer, etc.
Mixture control in the IDLE
CUT-OFF position.
Move control to the full RICH
position.
Engine flooded.
Refer to Pilot's Operating Handbook
Fuel selector valve in OFF
position. (Thru Serial
21064535).
Place selector valve in the ON
position to a cell known to contain gasoline.
Fuel ON-OFF valve in
OFF position (21064536
and on).
Place valve in ON position.
Correct cause of
12-5
MODEL 210 & T210 SERIES SERVICE MANUAL
12-13.
TROUBLE SHOOTING (Cont).
TROUBLE
ENGINE STARTS BUT
DIES, OR WILL NOT
IDLE.
PROBABLE CAUSE
REMEDY
Idle stop screw or idle mixture
incorrectly adjusted.
Refer to paragraph 12-46.
Spark plugs fouled or improperly
gapped.
Remove, clean and regap plugs.
Replace if defective.
Water in fuel system.
Open fuel strainer drain and check
for water. If water is present,
drain fuel tank sumps, lines and
strainer.
Defective ignition system.
Refer to paragraph 12-79.
Vaporized fuel. (Most likely to
occur in hot weather with a hot
Refer to Pilot's Operating Handbook
engine.)
ENGINE RUNS ROUGHLY,
WILL NOT ACCELERATE
PROPERLY, OR LACKS
POWER.
12-6
Induction air leaks.
Check visually.
cause of leaks.
Manual primer leaking.
Disconnect primer outlet line.
If fuel leaks through primer,
repair or replace primer.
Dirty screen in fuel control unit
or defective fuel control unit.
Check screen visually. Check
fuel flow through control unit.
Clean screen. Replace fuel control unit if defective.
Defective manifold valve or
clogged screen.
Check fuel flow through valve.
Replace if defective. Clean screen.
Defective engine-driven fuel
pump.
If engine continues to run with
electric pump turned on, but stops
when it is turned off, the enginedriven pump is defective. Replace
pump.
Defective engine.
Check compression. Listen for
unusual engine noises. Engine
repair is required.
Propeller control set in high
pitch position (low RPM).
Use low pitch (high RPM) position
for all ground operation.
Defective aircraft fuel system.
Refer to Section 13.
Restricted fuel injection lines
or discharge nozzles.
Check fuel flow through lines and
nozzles. Clean lines and nozzles.
Replace if defective.
Propeller control in high pitch
(low RPM) position.
Use low pitch (high RPM) for
all ground operations.
Correct the
Restriction in aircraft fuel
system.
Refer to Section 13.
Restriction in fuel injection
system.
Clean system. Replace any
defective units.
MODEL 210 & T210 SERIES SERVICE MANUAL
12-13. TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
REMEDY
MODEL 210 & T210 SERIES SERVICE MANUAL
haul, 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 with the lines and hoses being disconnected at
the firewall.
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.
conb.
b. Place
Place fuel
fuel selector
selector valve
valve on
on fuel
fuel ON-OFF
ON-OFF control in the OFF position.
c. Remove engine cowling in accordance with paragraph 12-3..
graph 12-3.
d. Disconnect battery cables and insulate terminals
as a safety precaution.
e.
Drain fuel
e. Drain
fuel stainer
strainer and
and lines
lines.
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
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.
WARNING |4.
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
12-8
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 all hot and cold air flexible ducts
and remove.
k. Remove exhaust system in accordance with paragraph 12-97.
1. 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
Disconnect oiltemperature wire at probe
Disconnect oil temperature wire at probe
below oil cooler.
5.
Disconnect
electricalwires
wire and
andwire
wire shieldshield5. Disconnect
electrical
ground
at alternator.
ground at alternator.
6. Disconnect exhaust gas temperature wires at
quick-disconnects.
7. Disconnect electrical wires at throttle microswitches.
8. Remove all clamps and lacings attaching
wires or cables to engine and pull wires and cables
aft to clear engine.
m. Disconnect lines and hoses as follows:
i. Disconnect vacuum hose at firewall.
2. Disconnect oil breather and vacuum system
oil separator vent lines where secured to the engine.
WARNING
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.
3. Disconnect fuel supply and vapor return hoses
at fuel pump.
Disconnect primer line at firewall fitting.
5. Disconnect fuel-flow gage hose at firewall.
6. Disconnect oil pressure line at firewall
fitting.
7. Disconnect manifold pressure hose at firewall.
8. Disconnect manifold and balance tube drain
lines.
n. 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.
MODEL 210 & T210 SERIES SERVICE MANUAL
CAUTION
~
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.
o. 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.
p. Remove bolts, ground strap and heat deflectors.
q. Slowly hoist engine out of nacelle and clear of
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.
r. Remove engine shock-mounts and ground strap.
NOTE
If shock-mounts will be reused, 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.
12-15. CLEANING. Clean engine in accordance
with instructions in Section 2.
12-16. ACCESSORIES REMOVAL. Removal of engine accessories for overhaul or for engine replacement involves stripping the engine of parts, accessories and components to reduce it to the bare engine. During the removal process, removed items
should be examined carefully and defective parts
should be tagged for repair or replacement with new
components.
through protective plys, cuts, breaks, stiffness,
connections. Excessive
heat on hoses will cause them to become brittle and
easily broken. Hoses and lines are most likely to
crack or break near the end fittings and support
points.
d. Inspect for color bleaching of the end fitting or
severe discoloration of the hoses.
~CAUTION~
damaged threads and loose
NOTE
Avoid excessive flexing and sharp bends
when examining hoses for stiffness.
e. Refer to Section 2 for replacement intervals for
flexible fluid carrying hoses in the engine compartment.
f. For major engine repairs, refer to the engine
manufacturer's overhaul and repair manual.
12-18. BUILDUP. Engine buildup consists of installatlon of parts, accessories and components to
the basic engine to build up an engine unit ready for
installation on the aircraft. All safety wire, lockwashers, nuts, gaskets and rubber connections
should be new parts.
12-19. 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.
NOTE
Items easily confused with similar items
should be tagged to provide a means of
identification when being installed on a
new engine. All openings exposed by the
removal of an item should be closed by
installing a suitable cover or cap over
the opening. This will prevent entry of
foreign material. If suitable covers are
not available, tape may be used to cover
the openings.
12-17. INSPECTION. For specific items to be inspected, refer to the engine manufacturer's manual.
a. Visually inspect the engine for loose nuts, bolts,
cracks and fin damage.
b. Inspect baffles, baffle seals and brackets for
cracks, deterioration and breakage.
c. Inspect all hoses for internal swelling, chafing
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.
NOTE
Be sure engine shock-mounts, spacers and
washers are in place as the engine is
lowered into position.
d. Install engine-to-mount bolts, then remove the
hoist and support stand placed under tail tie-down
fitting. Torque bolts to 300 +50 -0 lb-in.
e. Route throttle, mixture and propeller controls
to their respective units and connect. Secure controls in position with clamps.
Revision 3
12-9
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTES
8
REINFORCED MOUNTS CONTAIN MOULDED-IN WASHER
AT THIS LOCATION
ON ALL MODELS:
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Throughout the aircraft fuel system, from
the fuel bays 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), 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.
f. Connect lines and hoses as follows:
1. Connect manifold and balance tube drain
lines.
2. Connect manifold pressure hose at firewall.
3. Connect oil pressure line at firewall fitting.
4. Connect fuel-flow gage hose at firewall.
5. Connect primer line at firewall fitting.
6. Connect fuel supply and vapor return hose at
Pump.
7. Connect oil breather and vacuum system oil
separator vent lines where secured to the engine.
8. Connect vacuum hose at firewall.
9. Install clamps and lacings securing hoses and
lines to the engine to prevent chafing.
g. Connect wires and cables as follows:
1. Connect electrical wires and wire shielding
ground at alternator.
2. Connect cylinder head temperature wire at
probe.
CAUTION
When connecting 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.
3. Connect starter electrical cable at starter.
4. Connect tachometer drive shaft at adapter.
Be sure drive cable engages drive in adapter. Torque
housing attach nut to 100-lb. in.
.clearance
5. Connect exhaust gas temperature wires at
quick-disconnects.
6. Connect electrical wires at throttle microswitches,
7. Connect oil temperature wire to probe below
oil cooler.
8. Install clamps and lacings securing wires and
cables to engine, engine mount and brackets,
h. Install exhaust system in accordance with paragraph 12-97.
i. Connect all hot and cold air flexible ducts.
j. Install propeller and spinner in accordance with
instructions outlined in Section 14.
k. 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.
WARNING
Be sure magneto switch is in OFF position
when connecting switch wires to magnetos.
1. Clean and install nduction air filter in accordance with Section 2.
m. 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.
n. Check all switches are in the OFF position and
connect battery cables.
o. Rig engine controls in accordance with paragraphs 12-85, 12-86, 12-87 and 12-88.
p. Inspect engine installation for security, correct
routing of controls, lines, hoses and electrical wiring, proper safetying and tightness of all components.
q. Install engine cowling in accordance with paragraph 12-3.
r. Perform an engine run-up and make final adjustments on the engine controls.
12-20.
FLEXIBLE FLUID HOSES.
12-21. PRESSURE TEST. Refer to Section 2 for
pressure test intervals. Perform pressure test as
follows:
a. Place mixture control in the idle cut-off position.
b. Operate the auxiliary fuel pump in the high position.
c. Examine the exterior of hoses for evidence of
leakage or wetness.
d. Hoses found leaking should be replaced.
e. After pressure testing fuel hoses, allow sufficient time for excess fuel to drain overboard from
the engine manifold before attempting an engine start.
f. Refer to paragraph 12-17 for detailed inspection
procedures for flexible hoses.
12-22. REPLACEMENT.
a. Hoses should not be twisted on installation.
Pressure applied to a twisted hose may cause failure
or loosening of the nut.
b. Provide as large a bend radius as possible.
c. Hoses should have a minimum of one-half inch
from other lines, ducts, hoses or surrounding objects or be butterfly clamped to them.
d. Rubber hoses will take a permanent set during
extended use in service. Straightening a hose with
a bend having a permanent set will result in hose
cracking. Care should be taken during removal so
that hose is not bent excessively, and during reinstallation to assure hose is returned to its original
position.
e. Refer to Advisory Circular 43.13, Chapter 10,
for additional installation procedures for flexible fluid
hose assemblies.
Revision 2
12-11
MODEL 210 & T210 SERIES SERVICE MANUAL
12-23.
ENGINE BAFFLES.
12-24. DESCRIPTION. The sheet metal baffles installed on the engine direct the flow of air around the
cylinders and other engine components to provide
optimum cooling. Thee baffles incorporate rubberasbestos composition seals at points of contact with
the engine cowling and other engine components to
help confine and direct the airflow to the desired area.
It is very important to engine cooling that the baffles
and seals are in good condition and installed correctly.
The vertical seals must fold forward and the side
seals must fold upwards. Removal and installation of
the various baffle segments is possible with the cowling removed. Be sure that any new baffles seal properly.
12-25. CLEANING AND INSPECTION. The engine
baffles should be cleaned with a suitable solvent to
remove oil and dirt.
NOTE
The rubber-asbestos seal are oil and grease
resistant but should not be soaked in solvent
for long periods.
12-26. REMOVAL AND INSTALLATION. Removal
and installation of the various baffle segments is
possible with the cowling removed. Be sure that any
replaced baffles and seals are installed correctly and
that they seal to direct the airflow in the correct direction. Various lines, hoses, wires and controls
are routed through some baffles. Make sure that
these parts are reinstalled correctly after installation of baffles.
12-27. REPAIR. Repair of an individual segment of
engine baffle is generally impractical, since, due to
the small size and formed shape of the part, replacement is usually more economical. However, small
cracks may be stop-drilled and a reinforcing doubler
installed. Other repairs may be made as long as
strength and cooling requirements are met. Replace
sealing strips If they do not seal properly.
12-28.
ENGINE OIL SYSTEM.
12-29. DESCRIPTION. The oil system is of the full
pressure wet sump type. Refer to applicable engine
manufacturer's overhaul manual for specific details
and descriptions.
Inspect baffles for cracks in the metal and for loose
and/or torn seals. Repair or replace any defective
parts.
SHOP NOTES:
12-12
FIGURE 12-3 DELETED
MODEL 210 & T210 SERIES SERVICE MANUAL
12-30.
TROUBLE SHOOTING.
TROUBLE
NO OIL PRESSURE.
LOW OIL PRESSURE.
PROBABLE CAUSE
REMEDY
No oil in sump.
Check with dipstick. Fill sump
with proper grade and quantity
of oil. Refer to Section 2.
Oil pressure line broken,
disconnected or pinched.
Inspect pressure lines. Replace
or connect lines as required.
Oil pump defective.
Remove and inspect. Examine
engine. Metal particles from
damaged pump may have entered
engine oil passages.
Defective oil pressure gage.
Check with a known good gage.
If second reading is normal,
replace gage.
Oil congealed in gage line.
Disconnect line at engine and gage;
flush with kerosene. Pre-fill with
kerosene and install.
Relief valve defective.
Remove and check for dirty or defective parts. Clean and install;
replace valve if defective.
Low oil supply.
Check with dipstick. Fill sump
with proper grade and quantity
of oil. Refer to Section 2.
Low viscosity oil.
Drain sump and refill with proper
grade and quantity of oil.
Oil pressure relief valve spring
weak or broken.
Remove and inspect spring.
Replace weak or broken spring.
Defective oil pump.
Check oil temperature and oil
level. If temperature is higher
than normal and oil level is
correct, internal failure is evident. Remove and inspect.
Examine engine. Metal particles
from damaged pump may have
entered oil passages.
Secondary result of high oil
temperature.
Observe oil temperature gage for
high indication. Determine and
correct reason for high oil temperature.
Dirty oil screens.
Remove and clean oil screens.
12-13/12-14 Blank
MODEL 210 & T210 SERIES SERVICE MANUAL
12-30.
TROUBLE SHOOTING (Cont).
TROUBLE
HIGH OIL PRESSURE.
LOW OIL TEMPERATURE.
HIGH OIL TEMPERATURE.
PROBABLE CAUSE
REMEDY
High viscosity oil.
Drain sump and refill with proper
grade and quantity of oil.
Relief valve defective.
Remove and check for dirty or defective parts. Clean and install;
replace valve if defective.
Defective oil pressure gage.
Check with a known good gage. If
second reading is normal, replace
gage.
Defective oil temperature gage
or temperature bulb.
Check with a known good gage. If
second reading is normal, replace
gage. If reading is similar, the
temperature bulb is defective.
Oil cooler thermostatic
bypass valve defective
or stuck.
Remove valve and check for proper
operation. Replace valve if defective.
Oil cooler air passages clogged.
Inspect cooler core. Clean air
passages.
Oil cooler oil passages clogged.
Drain oil cooler and inspect
for sediment. Remove cooler
and flush thoroughly.
Thermostatic bypass valve
damaged or held open by
solid matter.
Feel front of cooler core with hand.
If core is cold, oil is bypassing
cooler. Remove and clean valve
and seat. If still inoperative, replace.
Low oil supply.
Check with dipstick. Fill sump
with proper grade and quantity
of oil. Refer to Section 2.
Oil viscosity too high.
Drain sump and refill with proper
grade and quantity of oil.
Prolonged high speed operation
on the ground.
Hold ground running above 1500
rpm to a minimum.
Defective oil temperature gage.
Check with a known good gage.
If second reading is normal.
Replace gage.
Defective oil temperature bulb.
Check for correct oil pressure, oil
level and cylinder head temperature. If they are correct, check
oil temperature gage for being defective; if similar reading is observed, bulb is defective. Replace bulb.
Revision 2
12-15
MODEL 210 & T210 SERIES SERVICE MANUAL
12-30.
TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
HIGH OIL TEMPERATURE
(Cont.)
REMEDY
Secondary effect of low oil
pressure.
Observe oil pressure gage for
low indication. Determine and
correct reason for low oil pressure.
Oil congealed in cooler.
This condition can occur only in
extremely cold temperatures.
If congealing is suspected, use
an external heater or a heated
hangar to warm the congealed oil.
OIL LEAK AT FRONT OF
ENGINE.
Damaged crankshaft seal.
Replace.
OIL LEAK AT PUSH ROD
HOUSING.
Damaged push rod housing oil seal.
Replace.
112-31.
FULL-FLOW OIL FILTER.
12-32. DESCRIPTION. An external oil filter may
be installed on the engine. The filter and filter
adapter replace the engine oil pressure screen. Beginning with the 1980 models a spin-on filter is used,
previous models used a replacement filter element
and filter can. The filter adapter incorporates a
bypass valve which will open allowing pressure oil
from the oil pump to flow to the engine oil passages
if the oil filter should become clogged on prior to
1980 models. The 1980 models have the bypass
valve in the spin-on oil filters.
12-33. REMOVAL AND INSTALLATION (FILTER
ELEMENT) (See figure 12-4).
NOTE
Filter element replacement kits and spin-on
filters are available from Cessna Parts
Distribution (CPD 2) through Cessna Service
Distribution (CPD 2) through Cessna Service
a. Remove engine cowling in accordance with
paragraph 12-3.
b. Remove both safety wires from filter can and
unscrew hollow stud (1) to detach filter assembly
from adapter (11) as a unit. Remove filter assembly from aircraft and discard gasket (9). Oil will
drain from filter as assembly is removed from
adapter.
c. Press downward on hollow stud (1) to remove
from filter element (5) and can (4). Discard metal
gasket (2) on stud (1).
d. Lift lid (7) off can (4) and discard lower gasket (6).
e. Pull filter element (5) out of can (4).
12-16
Revision 3
NOTE
knife,
Before discarding removed filter
element (5), remove the outer perforated paper cover; using a sharp
cut through the folds of the
filter element at both ends. Then,
carefully unfold the pleated element
and examine the material trapped in
the element for evidence of internal
engine damage, such as chips or
particles from bearings. In new or
newly overhauled engines, some small
particles or metallic shavings might
be found, these are generally of no
consequence and should not be confused with particles produced by impacting, abrasion or pressure.
Evidence of internal damage found in
the oil filter element justifies further
examination to determine the cause.
f. Wash lid (7), hollow stud (1), and can (4) in solvent |
and dry with compressed air.
NOTE
When installing a new filter element (5), it
is important that all gaskets are clean,
lubricated and positioned properly. Apply
a thin coating of Dow Corning compound,
DC-4, on the base gasket by brushing or
wiping. Also check that the correct amount
of torque is applied to the hollow stud (1).
If the stud is under-torqued. oil leakage will
occur. If the stud is over-torqued, the filter
can might possibly be deformed, again causing oil leakage.
MODEL 210 & T210 SERIES SERVICE MANUAL
14
^*..Ac-e
-
13
NOTE
Do NOT subsitute automotive gaskets for any
gaskets used in this assembly. Use only
approved gaskets listed in the Parts Catalogs.
12
d11
10
./
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
Hollow Stud
Metal Gasket)
Safety Wire Tab
Can
Filter Element
Lower Gasket
Lid
Thread Insert
Upper Gasket
Plug
Adapter
Bypass Valve
Nut (Adapter)
O-Ring
_____
40
SPIN - ON FILTER
BEGINNING WITH 21064136
_
30
2
THRU 21064135
T210 THRU 21064781
210 THRU 21064780
Figure 12-^.
Full-Flow Oil Filter
Revision 2
12-17
MODEL 210 & T210 SERIES SERVICE MANUAL
from can and cut off both ends. Carefully unfold the element and inspect for evidence of
internal engine damage such as chips or metal
from bearings. In new or newly overhauled
engines chips and bearing metal may be found.
and generally are of no consequence. However,
particles produced by impact, abrasion, or
pressure are evidence of internal engine damage
and justify further examination to determine the
cause.
12-33D. INSTALLATION.
a. Lightly lubricate filter gasket with engine oil or
Dow Corning Compound (DC-4).
b. Attach filter to adapter by turning clockwise until
it contacts base of adapter; then tighten 3/4 to one turn
or torque to 15 to 20 FT-LBS. Safety wire
c. Start engine and check for proper oil pressure;
w armup engine and check for filterper ol pesure;
d. Check that engine torque does not cause filter to
contact adjacent parts.
e. Replace engine cowl in accordance with paragraph 12-3.
f. Check oil level and filter leakage after operating
engine at high power setting, or after a flight around
the fielid.
12-34. FILTER ADAPTER. 210 THRU SERIAL
21064780, T210 THRU SERIAL 21064781.
NOTE
A special wrench adapter for adapter
nut (13) (Part No. SE-709) is available
from Cessna Parts Distribution (CPD 2)
through Cessna Service Stations, or one
may be fabricated as shown in figure
12-5. Remove any engine accessory that
interferes with removal of the adapter.
b. Note angular position of adapter (11), then remove
safety wire and loosen adapter nut (13).
c. Unscrew adapter and remove from engine. Discard
adapter O-ring (14).
12-36. DISASSEMBLY, INSPECTION AND REASSEM
BLY. Figure 12-4 shows the relative position of the
internal parts of the filter adapter and may be used
a a guide during installation of parts. The bypass
valve is to be installed as a complete unit, with the
valve being staked three places. The heli-coil type
insert (8) in the adapter may be replaced, although
special tools are required. Follow instructions of
the tool manufacturer for their use. Inspect threads
on adapter and engine for damage. Clean adapter in
solvent and dry with compressed air. Make sure all
passages in the adapter are open and free of dirt.
Check that bypass valve is seating properly.
12-35. REMOVAL. (Refer to figure 12-4.)
a. Remove filter assembly in accordance with paragraph 12-33.
Revision 3
12-18A/(12-18B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
VAPOR EJECTOR
TO TANK
PART
THROTTLE
POSITION
INTAKE AIR
FUEL INLET
FROM TANK
~
\ ~·
W^SS
\
VAPOR SEPARATOR
I
1
1
ETO
FUE
TOi\rf_
/
FLOW
MANIFOLD
GAG/
CALRAEVALVE
J
ADJUSTABLE ORIFICE
PUMP_
X
IT
\
RELIEF
.RELIEF
VALVE
--CHECK VALVE
|
-TO
^-VENT ___
FUEL INLET
RIEF
FUEL FLOW GAGE
FUESELD
PUMP
I/^
A
5
KMETERE\
S
AT
CALIBED
ORIFICE
RESSUREEN
P
T
A I
^MANIFOLD VALVE
AIRINLET
PRESSURE
i^
lNJECTION
MIXTURE
OUTLET
Detail
A
LEGEND:
I;;
I
E
RELIEF
VALVE PRESSURE
IMETERED FUEL
|
PUMP PRESSURE
^i
INLET PRESSURE
|
i
RETURN FUEL
Figure 12-6.
Fuel Injection Schematic
12-19
MODEL 210 & T210 SERIES SERVICE MANUAL
IDLE SPEED ADJUSTMENT
IDLE MIXTURE ADJUSTMENT
Figure 12-7.
Idle Speed and Idle Mixture Adjustment
on adapter and in engine for damage. Clean adapter
in solvent and dry with compressed air. Ascertain
that all passages in the adapter are open and free of
foreign material. Also, check that bypass valve is
seated properly.
12-37. INSTALLATION.
a. Assemble adapter nut (14) and new O-ring (15)
on adapter (11) in sequence illustrated in figure 12-4.
b. Lubricate O-ring on adapter with clean engine
oil. Tighten adapter nut until O-ring is centered in
its groove on the adapter.
c. Apply anti-seize compound sparingly to the adapter threads, then simultaneously screw adapter and
adapter nut into engie until .0-ring seats against engine boss without turning adapter nut (14). Rotate
adapter to approximate angular position noted during
removal. Do not tighten adapter nut at this time.
d. Temporarily install filter assembly on adapter,
and position so adequate clearance with adjacent parts
is attained. Maintainig this position of the adapter,
tighten adapter nut to 50-60 lb-ft (600-720 lb-in.) and
safety. Use a torque wrench, extension and adapter
as necessary when tightening adapter nut.
e. Using new gaskets, install filter assembly as
outlined in paragraph 12-33. Be sure to service the
engine oil system.
12-37A. FILTER ADAPTER. 210, BEGINNING
WITH 21064781; T210 BEGINNING WITH 21064782.
The oil filter adapter is an integral part of the oil
pump casting, located at the rear of the engine on
the right side.
12-38.
12-20
OIL COOLER.
Revison
12-39. DESCRIPTION. A non-congealing oil cooler
may be installed on the aircraft. Ram air passes
through the oil cooler and is discharged into the engine compartment. Oil circulating through the engine
is allowed to circulate continuously through warm-up
passages to prevent the oil from congealing when operating in low temperatures. On the standard and
non-congealing oilcoolers, as the oil increases to a
certain temperature, the thermostat valve closes,
causing the oil to be routed to all of the cooler passages for cooling. Oil returning to the engine from
the cooler is routed through the internally drilled oil
passages.
12-40. ENGINE FUEL SYSTEM.
12-6.)
(Refer to figure
12-41. 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 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.
MODEL 210 & T210 SERIES SERVICE MANUAL
FUEL METERING
ENGINE DRIVEN
UNIT
EXISTING FUEL PUMP
OUTLET HOSE
FUEL PUMP
NIPPLES
TEE
PRESSURE
INDICATOR
TEST HOSE
NIPPLE
TEST HOSE
NIPPLE
NOTE
WHEN ADJUSTING UNMETERED FUEL PRESSURE, TEST EQUIPMENT MAY
BE "TEED" INTO THE ENGINE-DRIVEN FUEL PUMP OUTLET HOSE AT THE
FUEL PUMP OR AT THE FUEL METERING UNIT.
Figure 12-8.
Fuel Injection Pump Adjustment Test Harness
NOTE
Throughout the aircraft fuel system, from
the fuel bays to the engine-driven pump,
use NS-40 (RAS-4) (Snap-On-Tools Corp.,
Kenosha, Wisconsin), MIL-T-5544 (Thread
Compound Antiseize, Graphite Petrolatum),
USP Petrolatum or engine oil as a thread
lubricator 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.
12-42.
FUEL-AIR CONTROL UNIT.
12-43. DESCRIPTION. This unit occupies the position ordinarily used for a carburetor, at the intake
manifold inlet. The function of this unit is to control
engine air intake and to set the metered fuel pressure
for proper fuel-air ratio. There are three control
elements in this unit, one for air and two for fuel.
One of the fuel control elements is for fuel mixture
and the other is for fuel metering. Fuel enters the
control unit through a strainer and passes to the
metering valve. The position of the metering valve
controls this fuel passed to the manifold valve and
nozzles. A linkage connecting the metering valve to
the air throttle proportions airflow to fuel flow.
The position of the mixture valve determines the
amount of fuel returned to the fuel pump. The fuel
control portion of the fuel-air control unit is enclosed in a shroud and is blast-air cooled to help
prevent vapor lock.
12-21
MODEL 210 & T210 SERIES SERVICE MANUAL
12-44. REMOVAL AND INSTALLATION.
a. Place all cabin switches and fuel selector or
fuel ON-OFF valve in the OFF position.
b. Remove cowling in accordance with paragraph
12-3.
c. Remove induction airbox in accordance with
paragraph 12-65.
d. Disconnect engine controls at throttle and mixture control arms.
NOTE
Cap all disconnected hoses, lines and fittings.
e. The three fuel lines which attach to the fuel control unit are routed inside flexible tubing to help cool
the fuel. Loosen tubing clamps at the control unit
and slide tubing back to gain access to the fuel line
fittings.
f. Disconnect fuel lines at control unit.
g. Loosen hose clamps which secure the control
unit to the right and left intake manifolds.
h. Remove control unit.
i. Cover the open ends of the intake manifold piping
to prevent entry of foreign matter.
j. Reverse the preceding steps for reinstallation.
Use new gaskets when installing control unit. Rig
throttle and mixture controls in accordance with paragraphs 12-85 and 12-86 respectively. Rig throttleoperated microswitch in accordance with Section 13.
12-45. CLEANING AND INSPECTION.
a. Check control connections, levers and linkage for
security, safetying and for lost motion due to wear.
b. Remove the fuel screen assembly and clean in
solvent (Stoddard or equivalent). Reinstall and safety.
c. Check the air control body for cracks and control unit for overall condition.
12-46. ADJUSTMENTS. (Refer to figure 12-7.) The
idle speed adjustment is a conventional spring-loaded
screw located in the air throttle lever. The idle
mixture adjustment is the locknut at the metering
valve end of the linkage. Tightening the nut to shorten the linkage provides a richer mixture. A leaner
mixture is obtained by backing off the nut to lengthen
the linkage. Idle speed and mixture adjustment
should be accomplished after the engine has been
warmed up. Since idle rpm may be affected by idle
mixture adjustment, it may be necessary to readjust
idle rpm after setting the idle mixture correctly.
a. Set the throttle stop screw to obtain 600 * 25
rpm, with throttle control pulled full out against idle
stop.
NOTE
Engine idle speed may vary among different
engines. An engine should idle smoothly,
without excessive vibration and the idle speed
should be high enough to maintain idling oil
pressure and to preclude any possibility of
engine stoppage in flight when the throttle is
closed.
12-22
b. Advance throttle to increase engine speed to
1000 rpm.
c. Pull mixture control knob slowly and steadily
toward the idle cut-off position, observing tachometer, then return control full IN (RICH) position
before engine stops.
d. Adjust mixture adjusting nut to obtain a slight
and momentary gain of 25 to 50 rpm at 1000 rpm
engine speed as mixture control is moved from full
IN (RICH) toward idle cut-off position. Return control to full IN (RICH) to prevent engine stoppage.
e. If mixture is set too LEAN, engine speed will
drop immediately, thus requiring a richer mixture.
Tighten adjusting nut (clockwise) for a richer mixture.
f. If mixture is set too RICH, engine speed will increase above 50 rpm, thus requiring a leaner mixture. Back off adjusting nut (counterclockwise) for
a leaner mixture.
NOTE
After each adjustment to the idle mixture,
run engine up to approximately 2000 rpm
to clear engine of excess fuel to obtain a
correct idle speed.
12-47. FUEL MANIFOLD VALVE (FUEL DISTRIBUTOR).
12-48. DESCRIPTION. Metered fuel flows to the
fuel manifold valve, which provides a central point
for distributing fuel to the individual cylinders. An
internal diaphragm, operated by fuel pressure,
raises or lowers a plunger to open and close the individual cylinder supply ports simultaneously. A
needle valve in the plunger ensures that the plunger
fully opens the outlet ports before fuel flow starts
and closes the ports simultaneously for positive
engine shut-down. A fine-mesh screen is included in the fuel manifold valve.
NOTE
The fuel manifold valves are supplied in two
flow ranges. When replacing a valve assembly, be sure the replacement valve has the
same suffix letter as the one stamped on the
cover of the valve removed.
12-49.
REMOVAL.
NOTE
Cap all disconnected lines, hoses and fittings.
a. Disconnect all fuel and fuel injection lines at
the fuel manifold.
b. Remove bolts which secure fuel manifold and
remove manifold.
12-50. CLEANING.
a. Remove manifold valve from engine in accordance with paragraph 12-49 and remove safety wire
from cover attaching screws.
MODEL 210 & T210 SERIES SERVICE MANUAL
cated in the cylinder heads. The outlet of each nozb. Hold the top cover down against internal spring
is directed into the intake port of each cylinder.
until all four cover attaching screws have been re- nozzle
The nozzle body contains a drilled central passage
moved, then gently lift off the cover. Use care not
with a counterbore at each end. The lower end is
to damage the spring-loaded diaphragm below cover.
used as a chamber for fuel-air mixture before the
c. Remove the upper spring and lift the diaphragm
spray leaves the nozzle. The upper bore contains an
assembly straight up.
orifice for calibrating the nozzles. Near the top,
radial holes connect the upper counterbore with the
NOTE
outside of the nozzle body for air admission. These
radial holes enter the counterbore above the orifice
If the valve attached to the diaphragm is
and draw outside air through a cylindrical screen
stuck in the bore of the body, grasp the
fitted over the nozzle body. This screen prevents
center nut, rotate and lift at the same
dirt and foreign material from entering the nozzle.
time to work gently out of the body.
A press-fit shield is mounted on the nozzle body and
over the greater part of the filter screen.
_extends
CAUTION
a small opening at the bottom of the shield.
~leaving
CAUTION
This provides an air bleed into the nozzle which aids
Do not attempt to remove needle or spring
in vaporizing the fuel by breaking the high vacuum in
from inside plunger valve. Removal of
the intake manifold at idle rpm and keeps the fuel
these items will disturb the calibration of
lines filled. The nozzles are calibrated in several
the valve.
ranges. All nozzles furnished for one engine are the
same range and are identified by a number and a
d. Using clean gasoline, flush out the chamber besuffix letter stamped on the flat portion of the nozzle
low the screen.
body. When replacing a fuel discharge nozzle be
e. Flush above the screen and inside the center
sure it is of the same calibrated range as the rest of
bore making sure that outlet passages are open. Use
the nozzles in the engine. When a complete set of
only a gentle stream of compressed air to remove
nozzles is being installed, the number must be the
dust and dirt and to dry.
same as the one removed, but the suffix letters may
be different, as long as they are the same for all
CAUTIONnozzles being installed on a particular engine.
The filter screen is a tight fit in the body and
12-54. REMOVAL.
may be damaged if removal is attempted. It
should be removed only if a new screen is to
NOTE
be installed.
f. Clean diaphragm, valve and top cover in the
same manner. Be sure the vent hole in the top cover
is open and clean.
g. Carefully replace diaphragm and valve. Check
that valve works freely in body bore.
h. Position diaphragm so that horizontal hole in
plunger valve is 90 degrees from the fuel inlet port
in the valve body.
i. Place upper spring in position on diaphragm.
i. Place cover in position so that vent hole in
cover is 90 degrees from inlet port in valve body.
Install cover attaching screws and tighten to 20±1
lb-in. Install safety wire on cover screws.
k. Install fuel manifold valve assembly on engine
in accordance with paragraph 12-51 and reconnect
all lines and hoses to valve.
1. Inspect installation and install cowling.
12-51. INSTALLATION.
a. Secure the fuel manifold to the crankcase with
the two crankcase bolts.
b. Connect the fuel lines and the six fuel injection
lines. Inspect completed installation and install
cowling.
12-52.
Plug or cap all disconnected lines and fittings.
a. Disconnect the fuel injection lines at the fuel discharge nozzles. Remove nozzles with a 1/2 inch deep
well socket wrench.
12-55. CLEANING AND INSPECTION. To clean
nozzles, immerse in clean solvent and use compressed air to dry them. When cleaning, direct air through
the nozzle in the direction opposite of normal fuel flow
Do not remove the nozzle shield or distort it in any
way. Do not use a wire or other metal object to clean
the orifice or metering jet. After cleaning, check the
shield height from the hex portion of the nozzle. The
bottom of the shield should be approximately 1/16
inch above the hex portion of the nozzle.
INSTALLATION.
12-56. INSTALLATION.
a. Install nozzles in the cylinders and tighten to a
torque value of 60 to 80 lb-in.
b. Connect the fuel lines at discharge nozzles.
c. Check installation for crimped lines, loose fittings, etc.
12-57.
FUEL INJECTION PUMP.
FUEL DISCHARGE NOZZLES.
12-53. DESCRIPTION. From the fuel manifold
valve, individual, identical size and length fuel lines
carry metered fuel to the fuel discharge nozzles lo-
12-58. DESCRIPTION. The fuel pump is a positivedisplacement, rotating vane type, connected to the
accessory drive section of the engine. Fuel enters
the pump at the swirl well of the pump vapor separa-
12-23
MODEL 210 & T210 SERIES SERVICE MANUAL
tor. 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 aircraft fuel
system. Since the pump is engine-driven, changes
inengine speed affects total pump flow proportionally. A check valve allows the auxiliary fuel pump
pressure to bypass the engine-driven fuel pump for
starting, or in the event of engine-driven fuel pump
failure. The pump supplies more fuel than is required by the engine; therefore, a spring-loaded,
diaphragm type relief valve is provided, with an adjustable orifice installed in the fuel passage to the
relief valve to maintain desired fuel pressure for
engine power setting. The adjustable orifice allows
the exact desired pressure setting at full throttle.
The fuel pump is equipped with a manual mixture
control to provide positive mixture control throughout
the range required by the injection system. This control limits output of the pump from full rich to idle
cut-off. Non-adjustable mechanical stops are located
at these positions. The fuel pump is ram-air cooled
to help prevent high fuel temperatures. The ram air
is picked up at the upper left engine baffle and directed through a flexible tube to the fuel pump shroud.
The fuel supply and return lines from the fuel pump
to the control unit are routed inside flexible tubes to
help prevent vaporized fuel at these points.
12-59. REMOVAL.
a. Place fuel selector or fuel ON-OFF valve in OFF
position and mixture control In IDLE CUT-OFF position.
b. Remove cowling in accordance with paragraph
12-3.
c. Loosen the clamps and slide the flexible tubes
free of the horns on the fuel pump shroud to gain access to the fuel lines.
d. Remove the alternator drive belt.
e. Tag and disconnect all lines and fittings attached to the fuel pump.
NOTE
Plug or cap all disconnected lines, hoses
and fittings.
f.
g.
Remove the shroud surrounding the fuel pump.
Remove the nuts and washers attaching the fuel
pump to the engine.
h.
Remove fuel pump and gasket.
WARNING
Residual fuel draining from lines and hose
constitutes a fire hazard. Use caution to
prevent accumulation of fuel when lines or
hoses are disconnected.
i. If a replacement pump is not being installed immediately, a temporary cover should be installed on
the fuel pump mount pad.
12-60. INSTALLATION.
a. Position a new gasket and fuel pump on the
mounting studs with fuel pump inlet to the left. Be
sure pump drive aligns with drive in the engine.
b. Secure pump to engine with plain washers, internal tooth lock washers and nuts. Tighten nuts
evenly.
c. Install cooling shroud on fuel pump
d. Install all fittings and connect all lines.
e. Install the flexible ram air tube on the air horn
of the fuel pump shroud and install clamp
f. Replace the alternator drive belt and tighten the
nuts on the adjusting arm so that the drive belt has
proper tension. Refer to Section 17.
g. Inspect completed installation.
12-61. ADJUSTMENT. The full rich performance
of the fuel injection system is controlled by manual
adjustment of the air throttle, fuel mixture and pump
pressure at idle and only by pump pressure at full
throttle. To make full rich adjustments, proceed as
follows:
a. Remove engine cowling in accordance with paragraph 12-3.
NOTE
Inspect the slot-headed adjustable orifice
needle valve (located just below the fuel
pump inlet fitting) to see if it is epoxy
sealed or safety wired to the brass nut.
Iftheneedle valve is epoxy sealed, Continental Aircraft Engine Service Bulletin
No 70-10 must be complied with before
calibration of the unit can be performed.
b. Disconnect the engine-driven fuel pump outlet
fitting or the fuel metering unit inlet fitting and "tee"
the test gage into the fuel injection system as illustrated in figure 12-8.
NOTE
Cessna Service Kit No. SK320-2J provides
a test gage, line and fittings for connecting
the test gage into the system to perform
accurate calibration of the engine-driven
fuel pump.
c. The test gage MUST be vented to atmosphere and
MUST be held as near to the level of the engine-driven
fuel pump as possible. Bleed air from test gage line
prior to taking readings.
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).
12-24
MODEL 210 & T210 SERIES SERVICE MANUAL
e. Adjust engine idle speed to 600 rpm and
check test gage for 9-11 PSI. Refer to figure 12-7
for idle mixture adjustment.
NOTE
Do not adjust idle mixture until idle pump
pressure is obtained.
DO NOT make fuel pump pressure adjustments while engine is operating.
f. If the pump pressure is not 9 to 11 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:
1. 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 12-12. If pressure is incorrect, stop engine
and adjust pressure by loosening locknut and turning
the slotheaded needle valve located just below the fuel
pump inlet fitting clockwise (CW) to increase pressure
and counterclockwise (CCW) 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 currect pressures are obtained, safety
adjustable orifice and orifice locknut.
1. Remove test equipment, run engine to check for
leaks and install cowling.
12-61A. AUXILIARY ELECTRIC FUEL PUMP
FLOW RATE ADJUSTMENT. Refer to Section 13.
INDUCTION AIR SYSTEM.
12-64.
AIRBOX.
12-65.
REMOVAL AND INSTALLATION.
a. Remove cowling in-accordance with paragraph
WARN I NG
12-62.
12-9.)
gine baffle. A spring-loaded alternate air door is
incorporated in the airbox and will open by engine
suction if the air filter should become clogged. This
permits unfiltered induction air to be drawn from
within the engine compartment.
(Refer to Figure
12-3.
b. Remove induction air filter.
c. Disconnect electrical wiring at throttle-operated
micro-switch and tape terminals as a safety "prec aution.
d. Remove clamps attaching lines, wires and controls to airbox.
e. Remove bolts securing airbox to fuel-air control
unit and engine and remove airbox and gasket.
f. Install a cover over fuel-air control opening.
g. Reverse the preceding steps for reinstallation.
Adjust throttle operated switch in accordance with
Section 13.
12-66. CLEANING AND INSPECTION. Clean metal
parts of the induction airbox with Stoddard solvent or
equivalent. Inspect for cracks, dents, loose rivets,
etc. Minor cracks may be stop-drilled. In case of
continued or severe cracking, replace airbox. Inspect alternate spring-loaded door for freedom of
operation and complete closing.
12-67.
INDUCTION AIR FILTER.
12-68. DESCRIPTION. An induction air filter,
mounted at the airbox inlet, removes dust particles
from the ram air entering the engine.
12-69. REMOVAL AND INSTALLATION.
a. Remove cowling in accordance with paragraph
12-3.
b. Remove bolts securing filter to the upper left
engine baffle and induction airbox inlet.
c. Reverse the preceding steps for reinstallation.
Make sure-the gasket is in place-between the filter
and airbox intake.
12-70. CLEANING AND INSPECTION. Clean and
inspect filter in accordance with instructions in Section 2.
12-71.
IGNITION SYSTEM.
(Refer to Figure 12-10.)
12-72. DESCRIPTION. The ignition system is comprised of two magnetos, two spark plugs in each cylinder, an ignition wiring harness, an ignition switch
mounted on the instrument panel and required wiring
between the ignition switch and magnetos.
12-63. DESCRIPTION. Ram air enters the induction air system through a filter at the upper left en-
12-25
MODEL 210 & T210 SERIES SERVICE MANUAL
12-73.
TROUBLE SHOOTING.
TROUBLE
ENGINE FAILS TO START.
ENGINE WILL NOT IDLE
OR RUN PROPERLY.
PROBABLE CAUSE
REMEDY
Defective ignition switch.
Check switch continuity.
if defective.
Replace
Spark plugs defective, improperly
gapped or fouled by moisture or
deposits.
Clean, regap and test plugs.
Replace if defective.
Defective ignition harness.
If no defects are found by a
visual inspection, check
with a harness tester. Replace defective parts.
Magneto "P" lead grounded.
Check continuity. "P" lead
should not be grounded in the
ON position, but should be
grounded in OFF position.
Repair or replace "P" lead.
Failure of impulse coupling.
Impulse coupling pawls should
engage at cranking speeds.
Listen for loud clicks as impulse couplings operate. Remove magnetos and determine
cause. Replace defective
magneto.
Defective magneto.
Refer to paragraph 12-79.
Broken drive gear.
Remove magneto and check magneto and engine gears. Replace
defective parts. Make sure no
pieces of damaged parts remain
in engine or engine disassembly
will be required.
Spark plugs defective, improperly gapped or fouled
by moisture or deposits.
Clean, regap and test plugs.
Replace if defective.
Defective ignition harness.
If no defects are found by a
visual inspection, check with
a harness tester. Replace
defective parts.
Defective magneto.
Refer to paragraph 12-79.
Impulse coupling pawls
remain engaged.
Listen for loud clicks as impulse
coupling operates. Remove
magneto and determine cause.
Replace defective magneto.
Spark plugs loose.
Check and install properly.
12-27/12-28 Blank
MODEL 210 & T210 SERIES SERVICE MANUAL
12-74.
MAGNETOS.
12-75. DESCRIPTION. The airplane may be
equipped with either 662 series or 6200 series
Slick magnetos. The magnetos contain a conventional two-pole rotating magnet (rotor),
mounted in ball bearings. Driven by the engine
through an-impulse coupling at one end, the rotor
shaft operates the breaker points at the other end
of the shaft. The nylon rotor gear drives a nylon
distributor gear which transfers high tension current from the wedge-mounted coil to the proper
outlet in the distributor block. A coaxial capacitor
is mounted in the distributor block housing to serve
as the condenser as well as a radio noise suppressor.
Both nylon gears are provided with timing marks for
clockwise or counterclockwise rotation. The distributor gear and distributor block having timing
marks, visible through the air vent holes, for
timing to the engine. A timing hole is located in
the 662 series magneto-in the bottom of the magneto
adjacent to the flange. In the 6200 series, the timing
hole is located in the distributor block. A timing pin
or 6-penny nail can be inserted through this timing
hole into the mating hole in the rotor shaft to lock
the magneto approximately in the proper firing position. The breaker assembly is accessible only after
removing the screws fastening the magneto halves
together and disconnecting the capacitor slip terminal.
Do not separate magneto halves while it is installed
on the engine.
12-76.
REMOVAL.
^~ 12-76.
REMOVAL.
a. Remove engine cowling in accordance with paragraph 12-3.
..
.timing
b. Tag for identification and remove high tension
holes.
wires from the magneto being removed.
WARNING
*
WARNING
The magneto is in a SWITCH ON condition
when the switch wire is disconnected. Remove the high tension wires from magneto
or disconnect spark plug leads from the
spark plugs to prevent accidental firing.
c. Disconnect switch wire from condenser terminal
at magneto. Tag wire for identification so it may be
installed correctly.
d. Rotate propeller in direction of normal rotation
until No. 1 cylinder is coming up on its compression
stroke.
NOTE
To facilitate the installation of a replacement
magneto, it is good practice to position the
crankshaft at the advanced firing angle for No.
1 cylinder during step "d." Any standard
timing device or method can be used, or if
the magneto being removed is correctly timed
to the engine, the crankshaft can be rotated to
a position at which the breaker points will be
just opening to fire No. 1 cylinder.
FIGURE 12-10 DELETED
e. Remove magneto retainer clamps, nuts and.
washers and pull magneto from crankcase mounting
pad.
NOTE
As the magneto is removed from its mounting be sure that the drive coupling rubber
bushing and retainer do not become dis-
lodged from the gear bub and fal into the
engne.
12-77. INTERNAL TIMING.
a. Whenever the gear on the rotor shaft or the cam
(which also serves as the key for the gear) has been
removed, be sure that the gear and cam are installed
so the timing mark on the gear aligns with the "0"
etched on the rotor shaft.
b. When replacing breaker assembly or adjusting
contact breaker points, place a timing pin (or 0. 093
inch 6-penny nail) through the timing hole into the
mating hole in the rotor shaft. Adjusting contact
breaker points so they are just starting to open in
this position will give the correct point setting. Temporarily assemble the magneto halves and capacitor
slip terminal and use a timing light to check that the
timing marks, visible through the ventilation plug
holes are approximately aligned.
NOTE
The side of the magneto with the manufacturers insignia has a red timing mark and
the side opposite to the insignia has a black
the side opposite to the insignia has a black
mark viewed through the vent plug
holes. The
The distributor
distributor gear
gear also
also has
has aa red
red
timing mark and a black timing mark.
These marks are used for reference only
when installing magneto on the engine.
Do
not place red and black lines together on
the same side.
c. Whenever the large distributor gear and rotor
gear have been disengaged, they must be engaged
with their timing.marks alignedfor correct-rotation.
Align the timing mark on the rotor gear with the
"RH" on the distributor gear. Care must be taken to
keep these two gears meshed in this position until the
magneto halves are assembled.
12-78. INSTALLATION AND TIMING TO ENGINE.
The magneto MUST be installed with its timing marks
correctly aligned, with the number one cylinder on
its compression stroke and with number one piston
at its advanced firing position. Refer to paragraph
12-12 for the advanced firing position of number one
piston
WARNING
The magneto is grounded through the ignition
switch, therefore, any time the switch (primary) wire is disconnected from the magneto,
the magneto is in a switch ON or HOT condition. Before turning the propeller by hand,
remove the high tension wires from the magRevision 2
12-29
MODEL 210 & T210 SERIES SERVICE MANUAL
neto or disconnect all spark plug leads to
prevent accidental firing of the engine.
To locate the compression stroke of number one cylinder, remove the lower spark plugs from each cylinder except number one cylinder. Remove the top
plug from number one cylinder. Place thumb of one
hand over the number one cylinder spark plug hole
and rotate the crankshaft in the direction of normal
rotation until the compression stroke is indicated by
positive pressure inside the cylinder lifting the thumb
off the spark plug hole. After the compression stroke
is obtained, locate number one piston at its advanced
firing position. Locating the advanced firing position
of number one cylinder may be obtained by use of a
timing disc and pointer, Timrite, protractor and
piston locating gage or external engine timing marks
alignment.
NOTE
External engine timing marks are located on
a bracket attached to the starter adapter,
with a timing mark on the alternator drive
pulley as the reference point.
In all cases, it must be definitely determined that the
number one cylinder is at the correct firing position
and on the compression stroke, when the crankshaft
is turned in its normal direction of rotation. After
the engine has been placed in the correct firing position, install and time the magneto to the engine in the
following manner.
NOTE
Install the magneto drive coupling retainer
and rubber bushings into the magneto drive
gear hub slot. Insert the two rubber bushings into the retainer with the chamfered
edges facing toward the front of the engine.
a. Turn the magneto shaft until the timing marks,
visible through the ventilation plug holes are aligned,
(red-to-red or black-to-black). Insert a timing pin
or . 093 inch diameter 6-penny nail through the timing
hole on the bottom of the magneto adjacent to the
flange (662 series); or in the distributor block (6200
series). Next, push the timing pin through the mating
hole in the rotor shaft. This locks the magneto close
to the firing position during installation on the engine.
then push it back into mesh. DO NOT WITHDRAW THE MAGNETO DRIVE GEAR FROM
ITS OIL SEAL.
b. After magneto gasket is in place, position the
magneto on the engine and secure, then remove the
timing pin from the magneto. Be sure to remove
this pin before turning the propeller.
c. Connect a timing light to the capacitor terminal
at the front of the magneto and to a good ground.
d. Turn propeller back a few degrees (opposite of
normal rotation) to close the contact points.
NOTE
Do not turn the propeller back far enough to
engage the impulse coupling or the propeller
will have to be turned in normal direction of
rotation until the impulse coupling releases,
then backed up to slightly before the firing
position.
e. Slowly advance the propeller in the normal direction of rotation until the timing light indicates the contact points breaking. Magneto mounting clamps may
be loosened so that the magneto may be shifted to
break the points at the correct firing position.
f. Tighten magneto mounting nuts and recheck
timing.
g. Repeat steps "a" through "f" for the other magneto.
h. After both magnetos have been timed, check synchronization of both magnetos. Magnetos must fire
at the same time.
i. Remove timing devices from magneto and engine.
j. Connect spark plug leads to their correct magneto outlets.
NOTE
The No. 1 magneto outlet is the one closest
to the ventilation plug on the side of the
magneto having the manufacturer's insignia.
The magneto fires at each successive outlet
in clockwise direction. Connect No. 1 magneto outlet to No. 1 cylinder spark plug lead,
No. 2 outlet to the next cylinder to fire, etc.
Engine firing order is listed in paragraph
12-12.
k. Connect ignition switch (primary) leads to the
NOTE
capacitor terminals on the magnetos.
1. Inspect magneto installation and install engine
cowling in accordance with paragraph 12-3.
If the magneto drive gear was disengaged
during magneto removal, hold the magneto
in the horizontal position it will occupy
when installed, make certain that the drive
gear coupling slot is aligned with the magneto coupling lugs. If it is not aligned, pull
the magneto drive gear out of mesh with its
drive gear and rotate it to the aligned angle,
12-79. MAINTENANCE. At the first 25-hour inspection and at each 100-hour inspection thereafter,
the breaker compartment should be inspected. Magneto-to-engine timing should be checked at the first
25-hour inspection, first 50-hour inspection, first
100-hour inspection and thereafter at each 100-hour
12-30
Revision
MODEL 210 & T210 SERIES SERVICE MANUAL
inspection. If timing is as specified in paragraph 1212, internal timing need not be checked. If timing is
out of tolerance, remove magneto and set internal
timing, then install and time to the engine. In the
event the magneto internal timing marks are off
more than plus or minus five degrees when the breaker points open to fire number one cylinder, remove
the magneto and check the magneto internal timing.
Whenever the magneto halves are separated the
breaker point assembly should always be checked.
As long as internal timing and magnet-t-toengine
timing are within the preceding tolerances, it is
recommended that the magneto be checked internally
only at 500 hour intervals. It is normal for contact
points to burn and the cam to wear a comparable
amount so the magneto will remain in time within12-80.
itself. This is accomplished by having a good area
making contact on the surface between the points
and the correct amount of spring pressure on the
cam. The area on the points should be twenty-five
percent of the area making contact. The spring
pressure at the cam should be 10. 5 to 12.5 ounces.
When the contact points burn, the area becomes
irregular, which is not detrimental to the operation
of the points unless metal transfer is too great which
will cause the engine to misfire. Figure 12-11 illustrates good and bad contact points. A small dent will
appear on the nylon insulator between the cam follower and the breaker bar. This is normal and does not
require replacement.
4. Check the carbon brush on the distributor
gear for excessive wear. The brush must extend a
minimum of 1/32 nch beyond the end of the gear
shaft. The spring which the carbon brush contacts
should be bent our approximately 20 degrees from
vertical, since spring pressure on the brush holds
the distributor gear shaft against the thrust bearing
in the distributor Mock.
5. Oil the bearings at each end of the distributor
gear shaft with a drop of SAE 20 oil. Wipe excess oil
from parts.
6. Make sure internal timing is correct and reassemble magneto. Install and properly time magneto to engine.
a. Moisture Check.
1. Remove magneto from engine and remove
screws securing the magneto halves together, disconnect capacitor slip terminal and remove distributor. Inspect for moisture.
2. Check distributor gear finger and carbon
MAGNETO CHECK. *'anced timing settings in some cases, is the r. :,t of the erroneous
practice of bumping magnetos up in timing in order
to reduce RPM drop on single igution. NEVER ADVANCE TIMING BEYOND SPECIFICATIONS IN ORDER TO REDUCE RPM DROP. Too much importance is being attached to RPM drop on single Ignition. RPM drop on single ignition is a natural characteristic of dual ignition design. The purpose of
the following magneto check is to determine that all
cylinders are firing. If all cylinders are not firing,
the engine will run extremely rough and cause for
investigation will be quite apparent. The amount of
RPM drop is not necessarily significant and will be
influenced by ambient air temperature, humidity,
airport altitude, etc. In fact, absence of RPM drop
should be cause for suspicion that the magneto timing
been bumped up and is set in advance of the setting
specified. Magneto checks should be performed on a
comparative basis between individual right and left
magneto performance.
a. Start and run engine until the oil and cylinder
head temperature is in the normal operating range.
b. Place the propeller control in the full low pitch
(high RPM) position.
c. Advance engine speed to 1700 RPM.
d. Turn the ignition switch to the "R" position and
note the RPM drop, then return the switch to the
"BOTH" position to clear the opposite set of plugs.
e. Turn the switch to the "L" position and note the
RPM drop, then return the switch to the "BOTH" posi-
brush for moisture.
tion.
NOTE
IfSi~~~~~ igniiontrobleshold eveophas
If ignition trouble should develop, spark plugs
and ignition wiring should be checked first. If
the trouble definitely is associated with a magneto, use the following to help disclose the
source of trouble without overhauling the magneto.
3. Check breaker point assembly for moisture,
especially on the surfaces of the breaker points.
4. If any moisture is evident in the preceding
places, wipe with a soft, dry, clean, lint-free cloth.
b. Breaker Compartment Check,
1. Check all parts of the breaker point assembly for security.
2. Check breaker point surface for evidence of
excessive wear, burning, deep pits and carbon deposits. Breaker points may be cleaned with a hardfinish paper. If breaker point assembly is defective,
install a new assembly. Make no attempt to stone or
dress the breaker points. Clean new breaker points
with clean, unleaded gasoline and hard-finish paper
before installing.
3. Check capacitor mounting bracket for cracks
or looseness.
. The RPM drop should not exceed 150 RPM on
either magneto or show greater than 50 RPM differentlal between magnetos. A smooth' RPM drop-off
past normal is uuallya sign of a too lean or too
rich mixture. A sharp RPM drop-off past normal is
usually a sign of a fouled plug, a defective harness
lead or a magneto out of time. If there is doubt concerning operation of the ignition system, RPM checks
at a leaner mixture setting or at higher engine speeds
will usually confirm whether a deficiency exists.
NOTE
An absence of RPM drop may be an indication of faultygrounding of one side of the
ignition system, a disconnected ground lead
at magneto or possibly the magneto timing
is set too far In advance.
12-31
MODEL 210 & T210 SERIES SERVICE MANUAL
THESE CONTACT
POINTS ARE USABLE
Figure 12-11.
CONTACT POINTS NEED
REPLACEMENT
Magneto Contact Breaker Points
12-81. SPARK PLUGS. Two spark plugs are installed in each cylinder and screw into hellcoil type
thread inserts. The spark plugs are shielded to prevent spark plug noise in the radios and have an internal resistor to provide longer terminal life. Spark
plug service life will vary with operating conditions.
A spark plug that is kept clean and properly gapped
will give better and longer service than one that is
allowed to collect lead deposits and is improperly
gapped.
NOTE
Refer to Section 2 for inspection intervals.
Remove, clean, inspect and regap all spark
plugs at these intervals. At this time, install
lower spark plugs in upper portion of cylinders and install upper spark plugs in lower
portion of cylinders. Since deterioration of
lower spark plugs is usually more rapid than
that of the upper spark plugs, rotating helps
prolong spark plug life.
12-82. ENGINE CONTROLS. (Refer to figure 12-11.)n
12-83. DESCRIPTION. The throttle, mixture and
propeller controls are of the push-pull type. The
propeller and mixture controls are equipped to lock
in any position desired. To move the control, the
spring-loaded button, located in the end of the control knob, must be depressed. When the button is
released, the control is locked. The propeller and
mixture controls also have a vernier adjustment.
Turning the control knob in either direction will
change the control setting. The vernier is primarily
for precision control setting. The throttle control
has neither a locking button nor a vernier adjustment,
but contains a knurled friction knob which is rotated
12-32
THESE
for more or less friction as desired. The friction
knob prevents vibration induced "creeping" of the control. A "Palnut" type locknut is installed in back of
the existing locknut at the engine end of the throttle,
mixture and propeller controls.
12-84. RIGGING. When adjusting any engine control,
it is important to check that the control slides smoothly throughout its full travel, that it locks securely if
equipped with a locking device and the arm or lever
which it operates moves through its full arc of travel.
CAUTION
Whenever engine controls are being disconnected, pay particular attention to the EXACT
position, size and number of attaching washers and spacers. Be sure to install attaching
parts as noted when connecting controls.
Refer to inspection and lubrication charts
in Section 2 of this manual for inspection,
lubrication and/or replacement intervals
for engine controls.
12-85. THROTTLE CONTROL.
a. Push throttle control full in, then pull control
out approximately 1/8 inch for cushion.
b. Check that throttle control arm is against the
mechanical stop. If necessary, loosen locknut and
screw rod end IN or OUT as necessary to align with
attachment hole while throttle arm is against the
mechanical stop.
c. Pull control full out and check that throttle arm
contacts the idle stop.
d. The throttle arm must contact the stops in each
direction and the control should have approximately
1/8 inch cushion when pushed full in.
MODEL 210 & T210 SERIES SERVICE MANUAL
12-86. MIXTURE CONTROL.
a. Push mixture control full in, then pull control
out approximately 1/8 inch for cushion.
b. Check that mixture control arm is in full rich
position (against stop). If necessary, loosen locknut
and screw rod end IN or OUT as necessary to align
with attachment hole while mixture arm is against
the mechanical stop.
c. Pull control full out and check that mixture arm
contacts the idle cut-off stop.
d. The mixture arm must contact the stops in each
direction and the control should have approximately
1/8 inch cushion when pushed full in.
NOTE
Refer to the inspection chart in Section 2 for
inspection and/or replacement interval for
the mixture control.
12-87. THROTTLE-OPERATED MICROSWITCH.
Refer to Section 13.
12-87A. LANDING GEAR WARNING HORN. Refer
to Section 5.
SHOP NOTES:
12-33
MODEL 210 & T210 SERIES SERVICE MANUAL
12-88.
14.
PROPELLER CONTROL.
Refer to Section
running clutch in the starter adapter, which incorporates worm reduction gears. The starter motor is
located just aft of the right rear cylinder.
12-89. STARTING SYSTEM.
12-90. DESCRIPTION. The automatically-engaged
starting system employs an electrical starter motor
mounted to a 90-degree adapter. A solenoid is activated by the ignition switch on the instrument panel.
When the solenoid is activated, its contacts close and
electrical current energizes the motor. Initial rotation of the motor engages the starter through an over-
12-91.
Never operate the starter motor more than
12 seconds at a time. Allow starter motor
to cool between cranking periods to avoid
overheating. Longer cranking periods
without cooling time will shorten the life
of the starter motor.
TROUBLE SHOOTING.
TROUBLE
STARTER WILL NOT OPERATE.
STARTER MOTOR RUNS, BUT
DOES NOT TURN CRANKSHAFT.
STARTER MOTOR DRAGS.
STARTER EXCESSIVELY
NOISY.
PROBABLE CAUSE
REMEDY
Defective master switch or circuit.
Check continuity.
switch or wires.
Install new
Defective starter switch or switch
circuit.
Check continuity.
switch or wires.
Install new
Defective starter motor.
Check electrical power to motor.
Repair or replace starter motor.
Defective overrunning clutch
or drive.
Check visually. Install new
starter adapter.
Starter motor shaft broken.
Check visually.
starter motor.
Low battery.
Check battery. Charge or
install new battery.
Starter switch or relay contacts
burned or dirty.
Install serviceable unit.
Defective starter motor
power cable.
Check visually.
cable.
Loose or dirty connections.
Remove, clean and tighten all
terminal connections.
Defective starter motor.
Check starter motor brushes,
brush spring tension, thrown
solder on brush cover. Repair
or install new starter motor.
Dirty or worn commutator.
Check visually. Clean and
turn commutator.
Worn starter pinion.
Remove and inspect.
starter drive.
Worn or broken teeth
on crankshaft gears.
Check visually. Replace
crankshaft gear.
Install new
Install new
Replace
12-35
MODEL 210 & T210 SERIES SERVICE MANUAL
12-92. PRIMARY MAINTENANCE. The starting
12-96A. ECONOMY MIXTURE INDICATOR (EGT)
circuit should be inspected at regular interals, the
Refer to Section 16.
frequency of which should be determined by the
12-97. REMOVAL AND INSTALLATION. (Refer to
amount of service and conditions under which the
figure 12-12. )
equipment is operated. Inspect the battery and wira. Remove engine cowling in accordance with paraing. Check battery for fully charged condition, prograph 12-3.
per electrolyte level with approved water and termib. Disconnect ducts from heater shroud on left mufassembly and EGT wires at quick-disconnects.
nals for cleanliness. Inspect wiring to be sure that fler
all connections are clean and tight and that the wiring
c. Disconnect tailpipe braces from shock-mounts at
insulation is sound. Check that the brushes slide
firewall brackets.
freely in their holders and make full contact on the
d. Remove nuts, springs and bolts attaching tailpipe
and muffler to collector pipe and remove muffler and
commutator. When brushes are worn to one-half of
tailpipe assemblies.
their original length, install new brushes (compare
brushes with new brushes). Check the commutator
e. Remove nuts attaching exhaust stack assemblies
for uneven wear, excessive glazing or evidence of
to the cylinders and remove exhaust stacks and gasexcessive arcing. If the commutator is only slightly
kets.
dirty, glazed or discolored, it may be cleaned with a
f. Reverse the preceding steps for reinstallation
strip of No. 00 or No. 000 sandpaper. If the commuInstall a new copper-asbestos gasket between each
riser and its mounting pad on each cylinder, regardtator is rough or worn, it should be turned in a lathe
less of apparent condition of those removed. Torque
and the mica undercut. Inspect the armature shaft
exhaust stack nuts at cylinders to 100-110 poundfor rough bearing surfaces. New brushes should be
inches.
by wrapping a strip
properly seated when installing
12-98. INSPECTION. Refer to Section 2 for inspecof No. 00 sandpaper around the commutator (with
tion intervals. Since exhaust systems of this type are
sanding side out) 1-1/4 to 1-1/2 times maximum.
subject to burning, cracking and general deterioration
Drop brushes on sandpaper covered commutator and
from alternate thermal stresses and vibrations, inturn armature slowly in the direction of normal respection is important and should be accomplished as
tation. Clean sanding dust from motor after sanding
specified in the Inspection Charts in Section 2. A
operations.
thorough inspection of the engine exhaust system is
required to detect cracks which could cause leaks
12-93. STARTER MOTOR.
and result in loss of engine power. To inspect the
12-94. REMOVAL AND INSTALLATION.
engine exhaust system, proceed as follows:
a. Remove engine cowling as required so that ALL
a. Remove engine cowling in accordance with parasurfaces of the exhaust assemblies can be visually
graph 12-3.
inspected.
NOTE
CAUTION
When disconnecting arter 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.
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. 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.
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.
12-95.
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 approxi-
EXHAUST SYSTEM.
12-96. 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.
12-36
NOTE
The inside of the vacuum cleaner hose should
be free of any contamination that might be
blown into the engine exhaust system.
MODEL 210 & T210 SERIES SERVICE MANUAL
1. Riser
2. Clamp Half
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
mately 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 air leak check as specified in step "b"
of this paragraph to make sure that the system is
acceptable.
e. In addition to the above inspections, at 200 hours
(after the mufflers have accumulated more than 1000
hours time in service) perform the following inspection:
1. Remove engine cowling in accordance with
paragraph 12-3.
2. Remove the mufflers from the collector
assemblies.
3. Remove the tailpipes from the mufflers.
4. Using a flashlight and a mirror, inspect
the baffles and cones from both ends of the mufflers.
Check for general deterioration and make sure the
baffles are intact and not separated from the support rods.
5. If defects are found, replace the mufflers
before further flight.
6. If no defects are found, reinstall the mufflers and tailpipes.
12-99.
CAUTIION
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 dislodging 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.
EXTREME WEATHER MAINTENANCE.
12-100. COLD WEATHER. Cold weather starting
will be made easier by the Installation of 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 forces fuel to the
engine. With an external power receptacle installed,
an external power source may be connected to assist
in cold weather or low battery starting. Refer to
paragraph 12-104 for use of the external power receptacle. The following may also be used to assist
engine starting in extremely cold weather. After
the last night 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 run-up after these conditions have
been followed, preheat the drained engine oil.
12-101. HOT WEATHER. Refer to Pilot's Operating
Handbook.
12-102. 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.
l
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
29ºC (-20-F), the engine compartment should be preheated by a ground heater. Pre-heating the engine
compartment is accomplished by inducing heated air
up through the cowl flap openings; thus heating both
the oil and cylinders. After the engine compartment
12-38
12-103. 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
filters daily as outlined in Section 2. Also change
engine oil and lubricate airframe items more often
than specified.
12-104. GROUND SERVICE RECEPTACLE.
to Section 17.
Refer
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 12A
ENGINE
TURBOCHARGED
WARNING
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the propeller,
since a loose or broken wire or a component
malfunction could cause the propeller to rotate.
NOTE
TABLE OF CONTENTS
For additional information covering turbocharger
and component maintenance, overhaul and trouble
shooting refer to the Manufacturer's Overhaul
Manual.
Page No.
Aerofiche/Manual
ENGINE COWLING ..........
2E7/12A-2
Description ...........
2E7/12A-2
Removal and Installation .....
2E7/12A-2
Cleaning and Inspection ......
2E7/12A-2
Repair
.............
2E7/12A-2
Cowl Flaps ...........
2E7/12A-2
Description
....
..
2E7/12A-2
Removal and Installation . . .2E7/12A-2
Rigging ...........
2E7/12A-2
ENGINE ..............
2E8/12A-3
Description .
.....
...
2E8/12A-3
Engine Data .
..........
2E8/12A-3
Time Between Overhaul (TBO) . . . 2E9/12A-4
Overspeed Limitations ......
2E9/12A-4
Trouble Shooting ........
2E10/12A-4A
Static Run-Up Procedures ....
.2E14/12A-8
Removal ..........
. .2E14/12A-8
Cleaning. ............
2E16/12A-10
Accessories Removal ......
. 2E16/12A-10
Inspection .
...........
2E16/12A-10
Buildup
............
2E16/12A-10
Installation ...........
2E16/12A-10
Flexible Fluid Hoses .....
. 2E17/12A-11
Pressure Test ........
2E17/12A-11
Replacement.........
2E17/12A-11
Engine Baffles ..........
2E17/12A-11
Description .
.......
2E17/12A-11
Cleaning and Inspection ....
2E17/12A-11
Removal and Installation
. . 2E17/12A-11
Repair
..........
.2E17/12A-11
ENGINE OIL SYSTEM .
......
2E18/12A-12
Description .
..........
2E18/12A-12
Trouble Shooting .
........
2E18/12A-12
Full-Flow Oil Filter .....
.2E18/12A-12
Description .......
. 2E18/12A-12
Removal and Installation .
. 2E18/12A-12
Filter Adapter .
.......
.2E18/12A-12
Removal
.....
2E18/12A-12
Disassembly, Inspection and
Reassembly .......
2E18/12A-12
Installation. .........
Oil Cooler
............
Description .........
ENGINE FUEL SYSTEM ......
Description ...........
Fuel-Air Control Unit ......
Description
.........
Removal
.
. ....
Cleaning and Inspection ...
Installation .
........
Adjustments
........
Fuel Manifold Valve .......
Description .........
Removal
.........
Cleaning
..........
Installation .........
Fuel Discharge Nozzles ......
Description
.........
Removal
..........
Cleaning and Inspection ..
Installation
Fuel Injection Pump .......
Description
........
Removal
.........
Installation .........
Adjustment (1977 thru 1982
Models) .
.........
Adjustment (Beginning with
1983 Models .......
INDUCTION AIR SYSTEM ......
Description
.
.....
Airbox
..
...........
Removal and Installation .
Cleaning and Inspection .
Induction Air Filter. ......
Description
.......
Removal and Installation .
Cleaning and Inspection ....
Installation of Induction Air
System Ducts .....
...
2E18/12A-12
2E18/12A-12
2E18/12A-12
.2E18/12A-12
2E18/12A-12
2E18/12A-12
2E18/12A-12
. 2E18/12A-12
2E20/12A-15
2E20/12A-15
.2E20/12A-15
2E20/12A-15
2E20/12A-15
2E20/12A-15
2E20/12A-15
2E20/12A-15
2E20/12A-15
2E20/12A-15
2E20/12A-15
.. 2E20/12A-15
. 2E20/12A-15
......
2E20/12A-15
2E20/12A-15
2E21/12A-16
2E21/12A-16
2E21/12A-16
.2E22/12A-16A
2E24/12A-17
.2E24/12A-17
2F1/12A-18
. 2F1/12A-18
. 2F1/12A-18
.2F1/12A-18
.2F1/12A-18
· . 2F1/12A-18
2F1/12A-18
2F2/12A-18A
Revision 3
12A-1
MODEL 210 & T210 SERIES SERVICE MANUAL
IGNITION SYSTEM
..........
2F2/12A-18A
Description ...........
2F2/12A-18A
Trouble Shooting .........
2F2/12A-18A
Magnetos .
...........
2F2/12A-18A
PressurizedMagnetos .....
2F2/12A-18A
Description .
........
2F2/12A-18A
Removal .........
2F2/12A-18A
Internal Timing .......
2F2/12A-18A
Installation and Timing-toEngine ..........
2F2/12A-18A
Maintenance .........
2F2/12A-18A
Magneto Check
........
2F2/12A-18A
Spark Plugs ...........
2F2/12A-18A
ENGINE CONTROLS .........
2F2/12A-18A
Description ...........
2F2/12A-18A
Rigging ............
2F2/12A-18A
Throttle Control .......
2F2/12A-18A
Mixture Control ......
.2F2/12A-18A
Propeller Control ......
2F2/12A-18A
Throttle Operated Microswitch. 2F3/12A-19
Auxiliary Electric Fuel Pump
Flow Adjustment ......
2F3/12A-19
Landing Gear Warning Horn . . 2F3/12A-19
STARTING SYSTEM .........
2F3/12A-19
Description ...........
2F3/12A-19
Trouble Shooting .........
2F3/12A-19
Primary Maintenance
.......
2F3/12A-19
Starter Motor ..........
2F3/12A-19
Removal and Installation . . . 2F3/12A-19
12A-1.
ENGINE COWLING.
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.
12A-4. CLEANING AND INSPECTION.
paragraph 12-4.
12A-5.
REPAIR.
12A-6.
COWL FLAPS.
Refer to
Refer to
Refer to paragraph 12-5.
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.
12A-8. REMOVAL AND INSTALLATION.
to paragraph 12-8.
Revision 3
b. Check to make sure that the flexible controls
reach their internal stops in each direction. Mark
controls so that full control travel can readily be
checked and maintained during the remaining rigging
procedures.
c. Place cowl flap control lever in the OPEN position, which is the top hole in the bracket. Be sure
that correct hole in bracket is used. If control lever
cannot be placed in correct hole in bracket, loosen
clamp at upper end of controls and slip housings in
clamp or adjust controls at upper clevis to position
control lever in correct hole in bracket.
d. THRU 1979 MODELS. Adjust clevis at lower end of
control so cowl flaps are streamlined in the closed
position. BEGINING WITH 1980 MODELS. Set cowl
open . 98 inch from cowl contour in the closed position.
Measure at outboard trailing edge of cowl flap and 90 °
to cowl skin. If full travel of the control is obtained
the open position will be correct.
f. Check that locknuts are tight. clamps are secure
and all bolts and nuts are installed.
Refer
12A-9. RIGGING.
a. Disconnect cowl flap control clevises from cowl
flaps.
12A-2
EXHAUST SYSTEM
........
. 2F3/12A-19
Description
..
......
... 2F3/12A-19
Removal ............
2F3/12A-19
Installation ...
.......
. 2F3/12A-19
Inspection ............
2F6/12A-22
TURBOCHARGER
..........
2F7/12A-23
Description ................
2F7/12A-23
Removal and Installation .....
2F7/12A-23
CONTROLLER AND WASTE-GATE
ACTUATOR ............
2F7/12A-23
Functions ............
2F7/12A-23
Operation ............
2F7/12A-23
Trouble Shooting .........
2F10/12A-26
Controller and Turbocharger Operational Flight Check. ......
2F14/12A-30
Removal and Installation of
Turbocharger Controller . . . . 2F15/12A-31
Absolute ControllerAdjustment . . . 2F15/12A-31
Removal and Installation of WasteGate and Actuator .......
2F15/12A-31
Adjustment of Waste-Gate
Actuator. ...........
2F16/12A-32
EXTREME WEATHER MAINTENANCE
2F16/12A-32
Cold Weather ..........
2F16/12A-32
Hot Weather .........
. 2F16/12A-32
Seacoast and Humid Areas . . . . 2F16/12A-32
Dusty Areas ...........
2F16/12A-32
Ground Service Receptacle ....
2F16/12A-32
NOTE
In all cases, the flexible controls must reach
their internal stops in each direction to assure
full travel of the controls.
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-10.
ENGINE.
12A-11. DESCRIPTION. An air-cooled, horizontally-opposed, direct-drive, fuel-injected, six-cylinder, turbocharged, Continental TSIO-520-R series
engine, driving a constant-speed propeller, is used
to power the aircraft The cylinders, numbered from
rear to 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
12A-12.
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 Supply Division.
ENGINE DATA.
Aircraft Series
T210
Model (Continental)
TSIO-520-R
BHP Maximum for Take-Off
(5 Minutes) at RPM
BHP Maximum Except Take-Off
RPM (maximum Continuous)
310
2700
285
2600
Limiting Manifold Pressure (Sea Level)
36.5 Inches Hg.
Number of Cylinders
6-Horizontally Opposed
Displacement
Bore
Stroke
520 Cubic Inches
5.25 Inches
4.00 Inches
Compression Ratio
7. 5:1
Magnetos
MagneSlick
Right Magneto
Left Magneto
Slick Model No. 662 (1977-1982 Models)
Model No. 6220 (Beginning with 1983 Models)
Fires 22 ° BTC Upper Right and Lower Left
Fires 22* BTC Upper Left and Lower Right
Firing Order
1-6-3-2-5-4
Spark Plugs
18mm (Refer to Continental Service Bulletin M77-10
for factory approved spark plugs and required gap)
33030 Lb-In.
Torque
Fuel Metering System
Unmetered Fuel Pressure
Nozzle Pressure
Continental Fuel Injection
5.5 to 6.5 PSI at 600 RPM
33 to 37 PSI at 2700 RPM (1977-1982 Models)
32 to 36 PSI at 2600 RPM (Beginning with 1983 Models)
3. 5 to 4.0 PSI at 600 RPM
19.0 to 20. 0 PSI at 2700 RPM
Oil Sump Capacity
With Filter Element Change
10 U.S. Quarts
11 U.S. Quarts
Tachometer
Mechanical Drive
Oil Pressure (PSI)
Minimum Idling
Normal
Maximum (Cold Oil Starting)
Connection Location
10
30-60
100
Between No. 2 and No. 4 Cylinders
Oil Temperature
Normal Operating
Maximum Permissible
Probe Location
Within Green Arc
Red Line (240°F)
In front of No. 5 Cylinder base
Revision 2
12A-3
MODEL 210 & T210 SERIES SERVICE MANUAL
Cylinder Head Temperature
Probe Location
Red Line (4600F) Max.
Lower Side No. 1 Cylinder (1977 thru 1979)
Without Airconditioning
With Airconditioning
No. 1
No. 3
No's. 1 or 5
No. 3
Lower Side of Cylinder
(1980 thru 1981)
(1982 and ON)
Economy Mixture Indicator (EGT)
Probe Location
Exhaust Collector R. H. Side (at turbine inlet)
Approximate Dry Weight With Accessories
(Excluding Turbocharger System)
461 Lb. (Weight is approximate and will vary
with optional accessories installed.)
12A-12A. TIME BETWEEN OVERHAUL (TBO).
Teledyne Continental Motors recommends engine
overhaul at 1400 hours operating time for the TSIO520-R series engines. Refer to Continental Aircraft
Engine Service Bulletin M79-14, Rev. 1, and to any
superseding bulletins, revisions or supplements
thereto, for further recommendations. 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
SHOP NOTES:
12A-4
FIGURE 12A-1 DELETEr
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-13.
TROUBLE SHOOTING.
TROUBLE
ENGLNE FAILS TO START.
ENGINE STARTS BUT DIES, OR
WILL NOT IDLE PROPERLY.
PROBABLE CAUSE
REMEDY
Engine flooded or improper use
of starting procedure.
Use proper starting procedure.
Refer to Pilot's Operating Handbook.
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
valve.
Remove and clean screen.
defective valve.
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-115.
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.
Replace
Replace
12A-4A/(12A-4B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
Clogged fuel screen in fuel manifold valve or defective valve.
Remove and clean. Replace
defective valve.
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-115.
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.
Replace
12A-5
MODEL 210 & T210 SERIES SERVICE MANUAL
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-100.
Turbocharger wheels rubbing.
Replace turbocharger.
Improperly adjusted or defective
waste-gate controller.
Refer to paragraph 12A-112.
Leak in turbocharger discharge
pressure system.
Correct cause of leaks. Repair
or replace damaged parts.
Manifold 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
MODEL 210 & T210 SERIES SERVICE MANUAL
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-112
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 Pilot's Operating Handbook.
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
MODEL 210 & T210 SERIES SERVICE MANUAL
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 Pilot's Operating Handbook
Defective engine.
Repair as required.
HIGH OR LOW OIL
TEMPERATURE
OR PRESSURE.
Refer to paragraph 12-30.
NOTE
Refer to paragraph 12A-107 for trouble shooting of controller
and waste-gate actuator.
12A-13A. STATIC RUN-UP PROCEDURES. In a
case of suspected lw engine power, a static runup should be conducted as follows:
a. Run-up engine, using take-off power and mixture 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 n atmospheric pressure,
temperature and humidity will have a
slight effect on static run-up.
c. Average the results of the RPM obtained. It
should be within 50 RPM of 2680 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 operation of alternate air door
spring or magnetic lock to make sure door will remain closed in normal operation.
12A-
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 correctunmetered fuel flow.
5. Check condition of induction air filter. Clean
if required.
6. Perform an engine compression check (Refer
to engine Manufacturer's Manual).
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 or fuel ON-OFF con-
trol 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
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
During the following procedures, remove
any clamps or lacings which secure controls, wires, hoses or lines to the engine,WARNING
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
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.
WARNING
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 1
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 microswitches.
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.
WARNIN
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
hoses at fuel pump.
pump drain line.
3. Disconnect
manifold.
4. Disconnect
wall.
5. Disconnect
fuel supply and vapor return
Disconnect and remove fuel
manifold pressure line at intake
the fuel-flow gage line at firethe 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.
1. Disconnect flexible ducting from heater shroud
and cabin valve.
m. 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.
CAUTION
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.
n. 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.
o. Remove mount bolts, ground strap and heat
shields.
p. Slowly hoist engine out of nacelle and clear of
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.
q. Remove engine shock-mounts.
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-9
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-15.
CLEANING.
Refer to paragraph 12-15.
12A-16. ACCESSORIES REMOVAL.
graph 12-16.
graph
NOTE
Refer to para-
Throughout the aircraft fuel system, from
the fuel bays to the engine-driven fuel
use NS-40 (RAS-4) (Snap-On Tools
12-16.pump,
12A-17.
INSPECTION.
12A-18.
BUILDUP. Refer to paragraph 12-18.
Refer to paragraph 12-17.
12A-19. INSTALLATION. Before installing the engine on the aircraft, instal any items which were
removed from the engine or aircraft after the engine
was removed.
NOTE
Remove all protective covers, plugs, caps
and identification tags as each item is connected or installed. Omit any items not
present on a particular
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.
NOTE
Be sure engine shock-mounts, spacers and
washers are in place as the engine is lowered
into position.
Corp.,
Wisconsin,
-T-5544
(Thread Kenosha,
Compound, Antiseize,
Graphite
Petrolatum),
USP Petrolatum
as
a thread lubricant
or to sealora engine
leakingoil
connection. Apply sparinglyto male threads
only,
cising
omitting the first to threads, exerextreme 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.
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.
3.
Connect the oil pressure line at its fitting.
Connect the fuel-flow gage line at firewall.
4. Connect the fuel supply and the vapor return
lies 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.
d. Attach ground strap under engine sump bolt and
7. Connect air and oil lines at waste-gate oninstall engine mount bolts. Torque bolts to 300 + 50 -0
troller on firewall.
lb-in. Bend tab washers to form lock for mount bolts.
8. Connect air vent line to fuel-flow gage line at
Install heat shields.
9. Connect engine primer lines at right and left
e. Remove support stand placed under tail tie-down
intake manifolds
fitting and remove hoist.
10. Connect oil drain line to oil deflector under
oil filter.
~external
~~~~~NOTE
11. Install all clamps securing lines and hoses to
If the exhaust system was loosened or reengine or structure.
moved, refer to paragraph 12A-99.
i. Connect wires and cables as follows:
1. Connect oil temperature wire at probe below
f. Connect flexible ducting on heater shroud and
oil cooler.
cabin valve.
2. Connect tachometer drive to adapter and torg. Route propeller governor control along left side
que to 100 lb-in.
of engine and secure with clamps.
SHOP NOTES:
12A-10
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
WARNING
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.
NOTE
When installing a new or newly overhauled
engine, and prior to starting the engine, tag
and 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 the filter connected,
operate the engine approximately 15 minutes
to filter out any foreign particles from the
oil. This is done to prevent foreign material
from entering the controller. After this run
period disconnect the full-flow filter and reconnect the lines to the controller as tagged.
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 wiring to throttle switches.
7. Connect exhaust gas temperature wires at
quick-disconnects.
8. Install clamps that attach wires or cables, to
r. Install engine cowling in accordance with paraengine or structure.
graph 12-3.
j. Connect engine controls and install block clamps.
s. Perform an engine run-up and make final adjustk. Rig engine controls in accordance with paraments on the engine controls.
graphs 12-85, 12-86, 12-87 and 12-88.
1. Install propeller and spinner in accordance with
12A-20. FLEXIBLE FLUID HOSES. Refer to parainstructions outlined in Section 14.
graph 12-20.
m. Complete a magneto switch ground-out and continuity check, then connect primary lead wires to the
12A-21 PRESSURE TEST. Refer to paragraph 12-21.
magnetos. Remove the temporary ground or connect
spark plug leads, whichever procedure was used dur-REPAIR
1
12A-22. REPLACEMENT. Refer to paragraph 12-22.
ing removal.
WARNING1c
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.
12A-23. ENGINE BAFFLES. Refer to paragraph
12-23.
12A-24. DESCRIPTION. Refer to paragraph 12-24.
12A-25. CLEANING AND INSPECTION.
paragraph 12-25.
Refer to
12A-26. REMOVAL AND INSTALLATION.
paragraph 12-26.
12A-27.
Refer to
REPAIR. Refer to paragraph 12-27.
SHOP NOTES:
12A-11
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-28.
ENGINE OIL SYSTEM,
12A-29.
DESCRIPTION.
The oil system
fold valve and the fuel discharge nozzles. The fuel
injection pump incorporates an adjustable aneroid
is of the full sensing unit which is pressurized from the discharge
pressure wet sump type. Refer to applicable engine
manufacturer's overhaul manual for specific details
and descriptions.
12A-30.
12-30.
TROUBLE SHOOTING.
discharge air pressure is also used to vent the fuel
discharge nozzles and the vent port of the fuel-flow
gage.
Refer to paragraph
NOTE
Throughout the aircraft fuel system, from the
fuel bays to the engine-driver fuel pump, use
NS-40 (RAS-4. Snap-On Tools Corp.. Kenosha,
Wisconsin), MIL-T-5544
Compound,
Wisconsin),
MIL-T-5544 (Thread
(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
12A-31. FULL-FLOW OIL FILTER. Referto paragraph 12-31.
12
11-32
12A-32. DESCRIPTION.
Refer to paragraph 12-32.
DESCRIPTION
Refer to paragraph 12-32.
12A-33. REMOVAL AND INSTALLATION.
to paragraph 12-33.
Refer
to
Refer
12A
ADAPTER
-34. FILTER
paragraph
12A-34. FILTER ADAPTER. Refer to paragraph
12A-35.
REMOVAL
Refer to paragraph 12-35.
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-36. DISASSEMBLY, INSPECTION AD REINSPECTION AND REASSEMBLY. Refer to paragraph
ASSEMBLY.
to paragraph
Refer
12-36.
12A-37.
INSTALLATION.
12A-38.
OIL COOLER.
12A-39.
DESCRIPTION.
12A-40.
ENGINE FUEL SYSTEM.
Refer to paragraph 12-37.
Refer to paragraph 138
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-fow type which controls fuel
flow to match engine airflow.
Any change in throttle
position, engine speed, or . combination of both,
causes changes in fuel flow in the correct relation to
causes engine
airflow. A manual mixture correct relation to
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 mani-
12A-12
12A-42. FUEL-AIR CONTROL UNIT.
paragraph 12-42.
12A-43.
12A-44.
DESCRIPTION.
Refer to
Refer to paragraph 12-43.
REMOVAL.
a. Place all cabin switches and fuel selector or
fuel ON-OFF valve in the OFF position.
fuel ON-OFF valve 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 elec-
tric fuel pump microswitch.
e. Disconnect throttle and mixture control rod ends
at fuel-air control unit.
NOTE
Cap or plug a
fittings.
disconnected hoses, lines and
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.
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
Refer to
paragraph 12-45.
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 wiring to throttle-operated
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.
12A-55. REMOVAL
a. Remove engine cowling in accordance with paragraph 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.
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.
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.
Refer to
FUEL INJECTION PUMP.
12A-59. DESCRIPTION. The fuel pump is a positive
displacement, rotating vane type. R 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 fight. 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
12A-15
MODEL 210 & T210 SERIES SERVICE MANUAL
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 forNOTE
fuel pump outlet pressure at idle.
12A-60. REMOVAL.
a. Place fuel selector or fuel ON-OFF valve in
OFF position.
b. Remove engine cowling in accordance with paragraph 12-3.
c. Remove alternator and left rear intake elbow.d.
d. Hoist engine far enough to remove weight from
engine mount and remove left rear engine mount leg,
shock-mount and alterntor 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 intallation. 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.DO
near to the level
heldasaspossible.Bleed
MUST
test gage
The
c. the
air
fuelbepump
driven
engine
of
from test gage line prior to taking readings.
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.
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 5.5 - 6.5 PSI. Refer to figure
12-7 for idle mixture adjustment.
NOTE
Do not adjust idle mixture until idle pump
pressure is obtained.
WARNING
NOT make fuel pump pressure adjustments while engine is operating.
12A-61. INSTALLATION.
f. If the pump pressure is not 5.5 - 6.5 PSI, stop
a. Install and align any fittings removed after pump
and turn the pump relief valve adjustment,
engine
removal.
pump clockwise (CW) to
fuel
the
on
aneroid
with
pump
install
gasket,
thin
b. Using new
increase pressure and counterclockwise (CCW) to
chamber down.
decrease pressure.
c. Install cooling shroud and remainder of fittings,
g. Mantaining idle pump pressure and idle RPM,
bending flanges of shroud to their original positions
correct idle mixture in accordance with paraobtain
removal.
during
noted
as
fittings
and aligning
12-46.
graph
d. Connect all fuel lines and shroud flexible duct.
h. Completion of the preceding steps have provided:
e. Install alternator bracket, shock-mount and
1. Correct idle pump pressure.
engine mount leg. Remove hoist, then adjust alterCorrect fuel flow.
2.
17.
Section
to
Refer
tension.
belt
drive
nator
Correct fuel metering cam to throttle plate
3.
f. Install intake elbow.
orientation.
g. Start engine and perform an operational check,
Advance to full throttle and maximum rated
adjusting fuel pump if required.
engine speed (propeller control full forward) with
h. Install cowling.
the mixture control in the full rich position and
12A-62. ADJUTMENT. (1977 thru 1982 Models).
pressure
manifold pressure
limit manifold
maximum limit
that maximum
verify that
Adjustments of the fuel injection pump requires
(36.5±. 5) is indicated. If manifold pressure is
special equipment and procedures. Adjustment to
at least
isis not
static RPM
or
RPM
2650 RPM
least 2650
not
the aneroid applies only to the fullthrottle setting. incorrect
RPM
or static
incorrect
or at
12A-110.
12A-13A
refer to paragraph
Adjustment of the idle position is obtained through
NOTE
the relief valve. To adjust the pump to the pressures
specified in paragraph 12A-12, proceed as follows:
If a static run-up, rated RPM (2700) cannot
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 connect
the test gage pressure bose and fittings into the fuel
Injection system as shown In figure 12A-3. Gage
MUST be vented to atmosphre.
NOT15E
Cessna
ervice KNit No. K320-2 K provides
Caessnagervice itNod fitigs for connecig
thetest age into the systemgs tor connecting
t ccutet calig tion of the esst nperformn
ncorrect,
eng -d
calibration
accurateof th
fuel pump
12A-16
throttle, adjust pump
be achieved at
limits (-1
each
flow
for each
PPH for
(-1 PPH
below limit
slightly below
flow slightly
pressures
correct
that
Verify
10 RPM low).
when rated
are
is achteved
achieved
RPM is
rated RPM
obtainedwhen
are obtained
during ta-off rol.
j. Check ships fuel flow gage for 186 - 190 PPH.
If fuel flow is Incorrect, stop engine and adjust flow.
This is accomplished by loosening the locknut and
turning the adjusting screw located at the rear of the
aneroid counterclockwise (CCW) to Increase flow or
clockwise (CW) to decrease flow. When fuel flow is
verify the unmetered pressure is within the
limits specified in paragraph 12A-12.
MODEL 210 & T210 SERIES SERVICE MANUAL
k.
After correct pressures are obtained, shut down
DO NOT make fuel pump pressure adjustments while engine s operating.
screw.
iscrew.
Remove test equipment, run engine to check for
leaks and install cowling.
12A-62A. ADJUSTMENT. (Beginning with 1983
Models. ) 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 or fuel limiter
unit and connect the test gage pressure hose and
fittings into the fuel injection system as shown in
figure 12A-3. Gage MUST be vented to atmosphere.
NOTE
Cessna Service Kit No. SK320-2K provides
a test gage, line and fittings for connecting
the test gage into the system to perform
accurate calibration of the engine-driven
fuel pump.
c. The test gage MUST be held as near to the level
of the engine driven fuel pump as possible. Bleed air
from test gage line prior to taking readings.
NOTE
The test gage should be checked for accuracy
at least every 90 days or anytime an error is
suspected. The tachometer iccuracy should
also be determined prior to making any adjustments to the pump.
d. Disconnect line from the return (center) port of
fuel flow limiter, plug line and cap port. See figure
12A-2A.
CAUTION
Do not plug side port (inlet) of pressure
limiter or limiter may be damaged during
adjustment.
l
g. If the pump pressure is not 5.5 - 6. 5 PSI, stop
engine and turn the pump relief valve adjustment,
on the centerline of the fuel pump clockwise (CW) to
increase pressure and counterclockwise (CCW) to
decrease pressure.
h. Maintaining idle pump pressure and idle RPM,
obtain correct idle mixture in accordance withparagraph 12-46.
. Completion of the preceding steps have provided:
1. Correct idle pump pressure.
2. Correct fuel flow.
3. Correct fuel metering cam to throttle plate
orientation.
J. Advance to full throttle and maximum rated
engine speed (propeller control full forward) with
the mixture control in the full rich position and
verify that maximum limit manifold pressure
(36. 5 . 5) is indicated. If manifold pressure is
incorrect or static RPM is not at least 2650 RPM
refer to paragraphs 12A-13A and 12A-110.
k. Retard the propeller control to obtain 2600 * 25
RPM stabilized.
Check ships fuel flow gage for 186 - 190 PPH.
If fuel flow is incorrect, stop engine and adjust flow
by loosening locknut and turning the adjusting screw
located at the aneroid counterclockwise
(CCW) to increase flow or clockwise (CW) to decrease
pressure is within the limits specified in paragraph
12A-12.
m. After correct pressures are obtained, shut down
engine and tighten locknut on fuel pump adjustment
screw.
n. Reconnect line to return (center) port of fuel
flow limiter.
o. Start engine and advance to full throttle with
mixture control full rich and the propeller control
full forward. Check the ships fuel flow gage for
186 - 190 PPH. If fuel flow is incorrect, shut down
the engine and adjust fuel flow set screw of. fuel flow
limiter (clockwise (CW) to increase, counterclockwise
(CCW) to decrease to obtain proper fuel flow.
p. Remove test equipment, run engine, check for
leaks and install cowling.
e. Start engine, warm up and run until oil temperature reads 40% to 70% in the green arc range. Oil
cooler inlet may have to be partially blocked in cold
weather. Set mixture control to full rich position and
propeller control full forward (low pitch, high RPM).
Revision 2
12A-16A
MODEL 210 & T210 SERIES SERVICE MANUAL
PRESSURE LIMITER
INSTALL CAP HERE
INSTALL CAP HERE
FUEL
METERING
UNIT
FUEL PUMP
Figure 12A-2A. Fuel-Injection Pump Adjustment/Test
12A-16B
Revision 2
MODEL 210 & T210 SERIES SERVICE MANUAL
FUEL METERING
UNIT
EXISTING FUEL PUMP
OUTLET HOSE
ENGINE DRIVEN
FUEL PUMP
NIPPLES
TEE
PRESSURE
INDICATOR
TEST HOSE
NIPPLE
TEST HOSE
NIPPLE
NOTE
WHEN ADJUSTING UNMETERED FUEL PRESSURE, TEST EQUIPMENT MAY
BE "TEED" INTO THE ENGINE-DRIVEN FUEL PUMP OUTLET HOSE AT THE
FUEL PUMP OR AT THE FUEL METERING UNIT.
Figure 12A-3.
12A-63.
Fuel Injection Pump Adjustment Test Harness (Turbocharged Engine)
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, Serial
21061574 thru 21063489, should be checked every 100
hours of operation for hinge wear, ease of operation,
and complete closing. If excessive hinge wear is
found, the hinge and magnetic catches should be replaced. Refer to Service Information Letter #SE80-12
for part numbers. The induction air filter should be
removed and cleaned at each 50-hour inspection; or
more frequently when operating under dusty conditions. Refer to Section 2.
Revision 2
12A-17
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
12A-70. REMOVAL AND INSTALLATION.
a. Remove right half of engine cowling in accordance with paragraph 12-3.
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.
NOTE
Clean filter and ascertain that induction air
ducts and airbox are clean when installing.
f.
12A-67. CLEANING AND INSPECTION.
paragraph 12-66.
12A-68.
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-18
Revision 2
Reverse the preceding steps for reinstallation.
Refer to
12A-71. CLEANING AND INSPECTION. Clean and
inspect filter in accordance with Section 2.
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-71A. INSTALLATION OF INDUCTION AIR
SYSTEM DUCTS. When cutting induction air system
ducts to length, the support wire should be cut back
far enough to bend back (Minimum bend radius, 1/8
inch) under the clamp and protrude 1/4 inch. Do not
break the bond between the wire and the fabric. Before tightening clamps, make sure there is no twist
or torque on the duct. If the duct is supported with
MIL-Y-1140 cord in place of wire, the preceding
installation applies except; MIL-Y-1140 cord has no
minimum bend radius requirements.
minimum
.
bend
The minimum installed bend radii for wire-supported
ducts in plane of bend, measured from the wall of
the duct, are as follows:
1. Neoprene - one ply, 1/4 diameter of the
maximum duct dimension.
2. Neoprene - two ply, and silicone - one ply.
1/3 diameter of the maximum duct dimension.
3. Silicone - two ply. 1/2 diameter of the maximum duct dimension.
NOTE
Ducts carrying filtered induction air may
not have local areas hand-formed to a
different cross section.
Refer to paragraph
12A-72.
12-71.
IGNITION SYSTEM.
12A-73.
DESCRIPTION.
12A-74.
12-73.
TROUBLE SHOOTING.
12A-75.
MAGNETOS.
Refer to paragraph 12-72.
Refer to paragraph
Refer to paragraph 12-74.
12A-75A. PRESSURIZED MAGNETOS (Beginning
with 1983 Model T210). Pressurized air is taken
from the throttle body adaptor assembly and directed
by a hose, through a filter, to a tee and then to each
magneto. The filter material is enclosed in a transparent case, with a flow arrow imprinted on it. The
filter should be replaced when the filtering material
is dirty.
Refer to paragraph 12-75.
12A-76.
DESCRIPTION.
12A-77.
REMOVAL.
12A-78.
12-77.
INTERNAL TIMING.
Refer to paragraph 12-76.
Refer to paragraph
12A-79. INSTALLATION AND TIMING-TO-ENGINE.
Refer to paragraph 12-78.
Refer to paragraph 12-79.
12A-80.
MAINTENANCE.
12A-81.
12-80.
MAGNETO CHECK.
12A-82.
SPARK PLUGS.
12A-83.
12-82.
ENGINE CONTROLS.
12A-84.
DESCRIPTION.
12A-85.
RIGGING.
12A-86.
12-85.
12A-87.
12-86.
12A-88.
14.
Refer to paragraph
Refer to paragraph 12-81.
Refer to paragraph
Refer to paragraph 12-83.
Refer to paragraph 12-84.
.
THROTTLE CONTROL. Refer to paragraph
MIXTURE CONTROL.
Refer to paragraph
PROPELLER CONTROL.
Revision 2
Refer to Section
12A-18A/(12A-18B blank
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-89. RIGGING THROTTLE-OPERATED MICROSWITCH. Refer to Section 13.
12A-98A. ELECTRIC
AUXILIARY
FUEL
12A-89A. AUXILIARY ELECTRIC FUEL PUMP
FLOW ADJUSTMENT. Refer to Section 13.
12A-89B. LANDING GEAR WARNING HORN.
Refer to Section 5.
12A-90. STARTING SYSTEM. Refer to paragraph
12-89.
12A-91.
DESCRIPTION.
12A-92.
12-91.
TROUBLE SHOOTING.
Refer to paragraph 12-90.
Refer to paragraph
12A-93. PRIMARY MAINTENANCE.
graph 12-92.
12A-94.
Refer to para-
STARTER MOTOR.
12A-95. REMOVAL AND INSTALLATION.haust
a. Remove cowling in accordance with paragraph
12-3.
b. Remove induction airbox in accordance with
paragraph 12A-66.
.
.k.
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-96.
EXHAUST SYSTEM.
Refer to figurerisers
12A-97. 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-98. 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 export 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.
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
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-99.
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-6). Tighten bolts.
12A-19
MODEL 210 & T210 SERIES SERVICE MANUAL
INTAKE
PIPE
ATTACHES
TO ENGINE
ATTACHES
TO CYLINDERS
HEAT
SHIELD
INTAKE
HERE
TAILPIPE
INSTALLED
ECONOMY MIXTURE
(EGT) PROBE
INSTALLED HERE
12A-20
MODEL 210 & T210 SERIES SERVICE MANUAL
4
1.
2.
3.
4.
5.
3
HEAT 14
DEFLECTORS
AND INSULATORS
Clamp
6.
Crankcase
7. Heat
Bolt Shield
Intake
Manifold
Balance
Tube
8.
Lockwasher
Heat Deflector
9.
Washer
Rivet
Figure 12A-4.
10.11. Insulation
Right Nosecap
12.
Skin
13. Retaining
Left Nosecap
14. Screw
Exhaust System (Sheet 2 of 2)
12A-21
MODEL 210 & T210 SERIES SERVICE MANUAL
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 safetied as re-
quired, then perform step "b" of paragraph 12A-100
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.
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 and remove
heater shroud 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
NOTE
The lower sections of turbocharger supports (index 8, figure 12A-6)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.
k. Install all parts but do not tighten attaching
clamps or bolts.
1. Torque nuts attaching risers to cylinders evenly
to 100-110 1b-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.
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 safetied as required. reinstall any parts removed for access, then
install nose caps and cowling.
12A-100. 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 50 hours of operation. Also, a thorough inspection of the engine exhaust system should
12A -22
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 pressure test
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 and the heat exchanger area may
be checked manually by feel, or by using a soap and
water solution and watching for jubbles. The exhaust
manifold in the heat exchanger area must be free of
air leaks. In other areas, forming of bubbles is acceptable; however, 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.
d. It is recommended that any components of the
exhaust system found defective be replaced before
the next flight.
e. After installation of exhaust system components,
recheck by performing the air pressure test to make
sure that system is acceptable.
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-101.
TURBOCHARGER.
12A-102. DESCRIPTION. The turbocharger is an
exhaust gas-driven compressor, or air pump, which
provides high velocity air to the engine intake manifold. The turbocharger is composed of a turbine
wheel, compressor wheel, turbine housing and compressor housing. The turbine, compressor wheel
and interconnecting drive shaft comprise one cornplete assembly and are the only moving parts in
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 induction intake manifold. Exhaust gas is then dumped
overboard through the exhaust outlet of the turbine
housing and exhaust tailpipe. Air is drawn into the
compressor through the induction air filter and 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 bypass the turbine.
12A-103. REMOVAL AND INSTALLATION. (Refer to
figure 12A-6).
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.
f. 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.
1. Reverse the preceding steps for reinstallation.
When installing the turbocharger, install a new gasket between exhaust manifold and turbine exhaust
inlet. Reinstall safety wire.
CAU TION
When installing cowling or turbine access
door, check that the clearence between
cowling or turbine access door and nose
gear doors is within presceibed limits of
.12 to .15 inches. Refer to SE77-15 for
details.
12A-104. CONTROLLER ANDWASTE-GATE
ACTUATOR.
12A-105. FUNCTIONS. The waste-gate actuator
and controller uses engine oil for power supply. The
turbocharger is controlled by the waste-gate, wastegate actuator, the absolute pressure and overboost
control valve. The waste-gate bypasses engine exhaust gas around the turbocharger turbine inlet.
The waste-gate actuator, which is physically connected to the waste-gate by mechanical linkage, controls the position of the waste-gate butterfly valve.
The absolute pressure controller controls the maximum turbocharger compressor discharge pressure,
the overboost control valve prevents an excessive
pressure increase from the turbocharger compressor.
12A-106. 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
k. Work turbocharger from aircraft through cowl
flap opening in lower cowling.
SHOP NOTES:
12A-23
MODEL 210 & T210 SERIES SERVICE MANUAL
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 17,000 feet, the
absolute pressure controller will continue to maintain
36.5 ±. 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.
CA OUTION
This turbocharged engine installation is equipped with a controller sustem which automatically
controls 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 maximums, 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 maximums.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 -25
MODEL 210 & T210 SERIES SERVICE MANUAL
13A-107. TROUBLE SHOOTING.
TROUBLE
UNABLE TO GET RATED
POWER BECAUSE MANIFOLD PRESSURE IS LOW.
PROBABLE CAUSE
REMEDY
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-110.
Replace controller if defective.
Defective actuator.
Refer to paragraph 12A-112. Replace 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.
ENGINE SURGES OR
SMOKES.
Defective controller.
Refer to paragraph 12A-110.
Replace if not adjustable.
Waste-gate actuator linkage
Refer to paragraph 12A-112.
binding.
Waste-gate actuator leaking
Replace actuator.
oil.
TURBOCHARGER NOISY
WITH PLENTY OF POWER
Turbocharger overspeeding from
defective or improperly adjusted
controller.
Refer to paragraph 12A-110.
Replace if defective.
Waste-gate sticking closed.
Correct cause of sticking. Refer
to paragraph 12A-110. Replace
defective parts.
Controller drain line (oil return
to engine sump) obstructed.
Clean line. Replace if defective.
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-112.
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-112.
Replaceif defective. Correct
cause of sluggish operation.
12A-26
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-107.
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-112.
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 INDICA-TOR
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-112.
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-27
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
12A-108. 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.
TAKE-OFF-ABSOLUTE CONTROLLER CHECK.
a. Cowl Flaps - Open.
b. Airspeed - 105 KIAS.
c. Oil Temperature - Middle of green arc.
d. Engine Speed - 2700 ± 25 RPM.
e. Fuel Flow - 192 LBS/HR * 6 LBS/HR (Full Rich Mixture).
f. Full Throttle M. P. - Absolute controller should maintain 36. 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 theinitial 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 - 105 KIAS.
c. Engine Speed - 2500 RPM.
d. Fuel Flow - Adjust mixture for 120.0 LBS/HR.
e. Part-Throttle M. P. - 30.0 in. Bg.
f. Climb to 17, 000 feet - Check part-throttle critical altitude during climb.
This part-throttle critical altitude is where manifold pressure startz 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 17, 000 feet, discontinue check and proceed to cruise check.
Outside Air Temperature
Part-Throttle Critical Altitude (80% Power)
Standard or Colder
20°F Above Standard
40°F Above Standard
Above 21,000 feet
13, 000 to 19, 000 feet
7, 000 to 13, 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-107). 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.
CRUISE - TURBOCHARGER PERFORMANCE CHECK.
a. Cowl Flaps - Closed.
b. Airspeed - Level flight.
c. Pressure Altitude - 17, 000 feet.
d. Engine Speed - 2700 RPM (5 minute limit).
e. Part-Throttle M.P. - 30.0 in. Hg.
f. Fuel Flow - Lean to 130.0 LBS/HR.
g. Propeller Control (1) Slowly decrease RPM until manifold pressure starts to drop, indicating waste gate is closed.
NOTE
If the waste gate closes at engine speeds lower than shown on the chart
in figure 12A-7, the turbocharger performance is normal. If the waste
gate closes at engine speeds higher than shown in figure 12A-7, refer
to the trouble shooting chart in paragraph 12A-107.
(2) Note outside air temperature and RPM as manifold pressure starts to drop, which should be
in accordance with the chart in figure 12A-7.
(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.
12A-30
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
ABOLUTE PRESSURE
SOUTEPRESUR
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-112. ADJUSTMENT OF WASTE-GATE ACTUATOR. (Refer to figure 12A-9.)
a. Remove waste-gate actuator in accordance with
paragraph 12A-111.
b. Plug actuator outlet port and apply a 50 to 60
psig air pressure to the inlet port of the actuator.
. Check for 0.00 inch gap between butterfly and waste-gate body as shown in figure 12A-9.
d. f 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
1A-9.
figure 12A-9.
as
in figure
shown in
as shown
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 0ring against cylinder wall.
. Remove air pressure line and plug from actuator.
. Install waste-gate and actuator as outlined in
paragraph 12A-111.
12A-32
FLAT-BLADED SCREWDRIVER
.
Figure 12A-8.
Controller
Controller Adjustment
12A-113. EXTREME WEATHER MAINTENANCE.
Refer to paragraph 12-99.
12A-114.
12-100.
COLD WEATHER.
12A-115. HOT
Handbook.
WEATHER.
Refer to paragraph
Refer to Pilot's Operating
12A-116. SEACOAST AND HUMID AREAS.
paragraph 12-102.
12A-117. DUSTY AREAS.
Refer to
Refer to paragraph 12-103.
12A-118. GROUND SERVICE RECEPTACLE.
to Section 17.
Refer
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 13
FUEL SYSTEM
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
FUEL SYSTEM. ........
.
...2F20/13-2
Description (THRU 21064535) .
. 2F20/13-2
Precautions ....
........
2F20/13-2
..
2F21/13-3
Trouble Shooting .....
Fuel Bays ........
.
. 2G2/13-8
Description
.........
2G2/13-8
Leaks .
..........
.. 2G2/13-8
Classification of Fuel Leaks . . 2G2/13-8
3s/qJ
-8
.
-Sealant .
......
. .. 2G3/13-9
Mixing .........
2G3/13-9
Sealing .....
. 2G3/13-9
Sealing Fuel Leaks ...
.2G3/13-9
Curing Time ...
2G6/13-12
Testing Fuel Bay
. ..
.2G6/13-12
Fuel Vents. ..........
.2G6/13-12
Description ........
.2G6/13-12
Removal and Installation
. . 2G7/13-13
Checking ..........
2G7/13-13
Fuel Quantity Indicating System . . 2G8/13-14
Description ........
.2G8/13-14
Removal and Installation . . . 2G8/13-14
Fuel Reservoirs (THRU 21064535) . 2G8/13-14
Description ........
2G8/13-14
Removal and Installation . . . 2G8/13-14
Fuel Selector Valve (THRU 21064535) .............
2G8/13-14
Description .
.........
2G8/13-14
Removal and Installation . . . 2G8/13-14
Repair .
...........
2G8/13-14
Auxiliary Fuel Pump
....
2G10/13-16
Description .........
2G10/13-16
Removal and Installation . . . 2G10/13-16
Circuit ..........
2G10/13-16
Rigging Throttle Operated
2G10/13-16
Microswitches .......
Flow Rate Adjustment ....
2G10/13-16
Maximum High Boost Check . . 2G12/13-18
Fuel Strainer (THRU 21064535).
. 2G12/13-18
Description ..........
2G12/13-18
Disassembly and Assembly . . 2G14/13-20
Removal and Installation . . .2G14/13-20
FUEL SYSTEM (BEGINNING WITH
2G14 /13-20
21064536) ..
...........
Description ........
.
2G1 4/13-20
Fuel Selector Valve
.......
2G14/13-20
.
....
. 2G14/13-20
.
Description
Removal and Installation . . .2G14/13-20
Disassembly, Repair and
........
2G14/13-20
Reassembly
.
Leak Test
.
.......
. 2G19/13-25
Alternate Method ....
2G19/13-25
Fuel Reservoir. .......
...2G19/13-25
Description .........
2G19/13-25
Removal and Installation ...
2G19/13-25
Fuel ON-OFF Valve ......
. 2G21/13-27
Description .........
2G21/13-27
Removal and Installation . . 2G21/13-27
Disassembly, Repair and
Reassembly .......
. 2G21/13-27
Fuel Strainer. . ........
.. 2G21/13-27
2G21/13-27
Description ........
Disassembly, Assembly and
2G21/13-27
Reassembly
........
Vented Fuel Filler Caps .....
2G24/13-30
Description ......
. 2G24/13-30
Metal "Flush-Type"
Filler Caps
........
2G24/13-30
. 2G24/13-30
Inspection .......
Cleaning .........
2G24/13-30
. 2G24/13-30
Reassembly .....
Red Plastic "Flush-Type"
.2G24/13-30
Filler Caps .......
2G24/13-30
Inspection
........
Cleaning
....
....
2H3/13-33
Reassembly .......
2H3/13-33
Leak Testing Metal or Red
2H3/13-33
.....
Plastic Filler Caps ..
13-1
MODEL 210 & T210 SERIES SERVICE MANUAL
13-1. FUEL SYSTEM.
The fuel system as defined
by this manual includes all components up to and including the fuel line connecting to the engine driven
pump inlet. Engine mounted components are covered
in Section 12 or 12A.
13-2.
DESCRIPTION.
(THRU 21064535.)
The fuel
system is essentially a gravity-flow system from the
bay outlets to the selector valve and a pump augmented
system from the selector valve to the engine. The
fuel system is comprised of the wing bays, reservoirs,
selector valve, auxiliary fuel pump, fuel strainer,
and associated plumbing. The fuel bay outlets are
located at the inboard end of the bays with lines subsequently routed down the front and rear doorposts,
under the floorboard, to the reservoirs. The fuel
line from the lower forward corner of each bay to the
reservoir serves as a combination fuel feed and vapor
return line. Fuel bypasses the auxiliary pump when
the pump is not in operation. The bays are individu-
ally vented overboard through vent lines with a check
valve located at each wing tip. Beginning with T210,
21063661 and earlier aircraft modified by SK210-93
the following changes have been made. The fuel lines
from the firewall to the strainer and the strainer to
the tunnel fitting will be changed from aluminum to
stainless steel with insulating sleeving. The fuel hose
from the fuel pump to the check valve and from the
check valve to the firewall and fuel pump to the tunnel
fitting will be changed from nonsleeved hose to fire
sleeved hose. The check valve is also fire sleeved.
SHOP NOTES:
13-2
13-3. PRECAUTIONS. During maintenance on the
fuel system the following precautions should be observed:
a. Aircraft should be properly GROUNDED prior to
performing maintenance on the fuel system or components.
b. Drain all lines or hoses when disconnected, because residual fuel draining constitutes a fire hazard,
and accumulation of this drainage increases the haz-
ard.
c. Cap open lines and cover connections to prevent
entry of foreign material in the former case, and
prevent damage to threads in the latter.
NOTE
Use NS-40(RAS-4) (Snap-On-Tools Corp.,
Kenosha, Wisconsin), 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, omit-
ting first two to prevent entry into fuel system. Use only a fuel soluble lubricant on
fitting threads, and use NO compound on the
injection system.
MODEL 210 & T210 SERIES SERVICE MANUAL
13-4.
TROUBLE SHOOTING.
Use this trouble shooting chart in conjunction with the engine trouble shooting chart in Section 12 or 12A.
TROUBLE
NO FLOW TO
ENGINE-DRIVEN
FUEL PUMP.
FUEL STARVATION AFTER
STARTING.
NO FUEL FLOW WHEN
ELECTRIC PUMP
OPERATED.
NO FUEL QUANTITY
INDICATION.
FLUCTUATING FUEL
PRESSURE INDICATIONS. (T210).
PROBABLE CAUSE
REMEDY
Fuel selector or fuel ON-OFF valve
not turned on.
Fuel bays empty.
Turn selector or fuel ON-OFF
valve on.
Service with proper grade and
amount of fuel.
Fuel line disconnected or broken.
Connect or repair fuel lines.
Fuel bay outlet screens plugged.
Remove and clean screens and
flush out fuel bays.
Defective fuel selector valve.
Repair or replace selector valve.
Plugged fuel strainer.
Remove and clean strainer and screen.
Defective check valve in electric
fuel pump.
Repair or replace pump.
Fuel line plugged.
Clean or replace fuel line.
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.
Plugged fuel vent.
Refer to paragraph 13-19.
Water in fuel.
Drain fuel bays, lines and
strainer.
Defective fuel pump switch.
Replace defective switch.
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 bays empty.
Service with proper grade and
amount of fuel.
Open or defective circuit breaker.
Reset.
Loose connections or open
circuit.
Tighten connections; repair or
replace wiring.
Defective fuel quantity indicator or transmitter.
Refer to Section 16.
Obstructed filter in fuel inlet
strainer of metering unit.
Remove and clean.
Manifold valve.
Replace.
Fuel flow indicator.
Replace.
Replace if defective.
13-3
MODEL 210 & T210 SERIES SERVICE MANUAL
MODIFIED PER CESSNA SERVICE KIT SK210-
8. Union
20.
Check Valve
MODEL 210 & T210 SERIES SERVICE MANUAL
*
SEE FIGURE 13-5
7
NOTE
21064102 thru 21064535 torque 340 to
380 lb-in. Lubricate threads per
MIL-H-5606.
8
Detail
21064136 thru
21064535
17
D
•24
24*
SEE FIGURE 13-6
WHEN MODIFIED BY SK210-138, AN
ADDITONAL FUEL DRAIN VALVE IS
INSTALLED IN THE OUTBOARD END
OF EACH FUEL TANK.
19
13
15
18
SEE FIGURE 13-7
23
Detail E
Figure 13-2.
13-6
Revision 3
Detail F
Fuel System (Sheet 2 of 2)
FUEL SAMPLER CUP
For use with drain
valves. (Refer to Section 2 of this manuaL)
MODEL 210 & T210 SERIES SERVICE MANUAL
The following procedure may be used to purge the
bay with argon or carbon dioxide.
a. Ground the aircraft to a suitable ground stake.
b. Set fuel selector valve handle in "OFF" posttion.
c. Drain all fuel from bay being repaired. (Observe
the precautions in paragraph 13-3.)
d. Remove access doors and insert hose to each
end of bay simultaneously.
e. Allow inert gas to flow into bay for several minutes (time dependent upon hose size, rate of flow,
etc.) to remove all fuel vapors.
Since argon or carbon dioxide are heavier than air,
these gases will remain in the bay during the repair.
The repair shall be made using non-sparking tools
(air motors, plastic scrapers, etc.)
completely around the joint when the parts are riveted
or fastened together. The fillet seal is applied after
the joint is fay surface sealed and riveted or fastened
together. Fillet sealing is applying sealant to the
edge of all riveted joints, joggles, bend reliefs, voids.
rivets or fasteners through the boundary of the bay
and any place that could produce a fuel leak. The fay
sealant need not be cured before the fillet seal is
applied, but the squeezed out sealant, to which the
fillet sealant is applied, must be free of dirt and contamination. Fillets laid on intersecting joints shall be
joined together to produce a continuous fillet.
Filler
sealant must be pressed into the joint, working out
all entrapped air. The best method of applying sealant is with an extrusion gun. Then work the sealant
into the joint with a small paddle, being careful to
eliminate all air bubbles.
NOTE
NOTE
Portable vapor detectors are available to
determine presence of explosive mixtures
and are calibrated for leaded fuel. These
detectors can be used to determine when
it is safe to make repairs.
13-10. FUEL BAY SEALANT. Two typesealants
are used in integral fuel bay construction. A pliable
type for access doors, and the rigid type for sealing
ribs and spars to the skin. Service Kit SK210-56C, seal
available through Cessna Supply Division, contains
these sealants with the proper ratio of acceleratorsNOTE
for each.
Keep sealants away from heat and flame.
Use only in a well ventilated area. Avoid
skin and eye contact. WEAR EYE SHIELDS.
In case of eye contact, flush generously
with clean water, and secure prompt medical attention.
13-12.
SEALING.
procedures).
Mix sealant according to
(Refer to Section 18 for repair
CAUTION
Protect
screens when
Protect drains
drains and
and fuel
fuel outlet
outlet screens
when
applying sealants to fuel bays.
Any repair that breaks the fuel bay seal will necessi-
tate resealing of that area of the bay.
a. Removeall existing sealant from area to be
sealed, leaving a taper on the remaining sealant.
The
taper will allow a scarf bond and a
Thetaperwillallowascarfbondand
a continuous
continuous
when the new sealant is applied.
The best method for removing sealant is
with a chisel tool made of hard fiber.
Remaining sealant is then removed with
WARNING
13-11. MIXING SEALANT.
service kit instructions.
During structural repair, parts must be predrilled, countersunk or dimpled and cleaned
before being sealed and positioned for final
installation.
Repair parts
that need sealing must beinstalled and riveted during
the sealing operation. All joints within the boundary
of the bay, but which do not provide a direct fuel path
out of the bay, such as stringers and rib flanges within the bay, must be fay surface sealed only. Joints
which provide a direct fuel path out of the bay area,
such as fuel spar flanges and inboard and outboard
rib flanges, must be fay surface sealed and fillet
sealed on the fuel side. Fay surface sealing is ap.
assembly
plying sealant to one mating part before assembly.
Enough sealant must be applied so it will squeeze out
aluminum wool. Neither steelwool nor
sandpaper can be used.
b. Vacuum thoroughly to remove all chips, filings,
and other foreign material from bay areas.
c. All surfaces and areas to be sealed shall be
thoroughly cleaned by wiping with a clean cloth
dampened with Methyl Ethyl Ketone (MEK), acetone
or similar solvent, and dried with a clean cloth prior
to solvent evaporation. Always pour the solvent on
the cloth. Never use contaminated solvent. The
cloth shall not be so saturated that dripping occurs.
13-13. SEALING FUEL LEAKS. First determine
the source of the fuel leak. Fuel can flow along a
seam or structure of the wing for several inches,
making the leak source difficult to find. A stained
area is an indication of the leak source. Fuel leaks
can be found by testing the complete bay as described
in paragraph 13-15. Another method of detecting the
source of a fuel leak is to remove access doors and
blow with an air nozzle from the inside of the bay in
the area of the leak while soap bubble solution is applied to the outside of the bay. After the leak source
has been found, proceed as follows:
a. Remove existing sealant in the area of the leak.
b. Clean the area and apply a fillet seal. Press
sealant into leaking area with a small paddle, working out all air bubbles
c. If leakage occurs around a rivet or bolt, restrike
c. I f
l e a k age oc cur sar ound
a r i v et or bol t
r
est r i ke
the rivet or loosen bolt, retorque, and reseal around
nutplate.
Revision 2
13-9
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Refer to paragraph 13-12.
TYPICAL
MODEL 210 & T210 SERIES SERVICE MANUAL
d. Apply fay surface door sealant to access doors,
fuel quantity transmitters, etc., if removed, and
install
e. Test fuel bay for leakage.
13-14. CURING TIME. Service Kit SK210-56 contains SP654706B2 Access Door Sealant Kit and
SP65489ºB2 Fuel Bay Sealant Kit. Normal curing
time for each seal is 24 hours. These values are
based on a standard condition of 77ºF (25°C) and
50% relative humidity. Curing time may be accelerated as shown in the following chart.
NOTE
hours. Curing time may be accelerated by applying
heat up to 120ºF on the PR1321B 1/2, and by applying
heat up to 130ºF on the PR1422B 1/2. Refer to
Accelerated Curing Time Chart above.
13-15. TESTING INTEGRAL FUEL BAY.
a. Remove vent line from vent fitting and cap fitting.
b. Disconnect fuel lines from bay.
c. To one of the bay fittings. attach a water manometer capable of measuring twenty inches of water.
d. To the other bay fitting, connect a well regulated
supply of air (1/2 PSI MAXIMUM, or 13. 8 INCHES of
water). Nitrogen may be used where the bay might
be exposed to temperature changes while testing.
e. Make sure filler cap is installed and sealed.
CAUTION
Temperature shall not exceed 160°F (71C).
Bay must be vented to relieve pressure
during accelerated curing.
Do not attempt to apply pressure to the bay
without a good regulator and a positive shutoff in the supply line. Do not inflate the fuel
bay to more than 1/2 psi or damage may occur.
ACCELERATED CURING TIME
*F of Sealant
160
140
130
120
Time in Hours
3
4
5 1/2
7
Service Kit SK210-101 contains PR1321B 1/2 Access
Cover Sealant Kit and PR1422B 1/2 Fuel Bay Sealant
Kit. Normal curing time for PR1321B 1/2 seal based
on a standard condition of 75-F (23. 9C) and 60%
relative humidity is 18 hours. Normal curing time
to
for PR1422B 1/2 seal based on a standard condition
of 75°F (23. 93º and 50% relative humidity is 45
SHOP NOTES:
13-12
Apply pressure slowly until 1/2 PSI is obtained.
Apply soap solution as required.
Allow 15 to 30 minutes for pressure to stabilize.
i.
If bay holds for 15 minutes, without pressure
loss, bay is acceptable.
j. Reseal and retest if any leaks are found.
f.
g.
h.
13-16.
FUEL VENTS.
13-17. DESCRIPTION. The fuel bay vent line extends
from the upper aft outbrd corner of each fuel bay
to the wing tip. This vent line contains a check valve
prevent fuel drainage through the vent line, but
MODEL 210 & T210 SERIES SERVICE MANUAL
13-20. FUEL QUANTITY INDICATING SYSTEM.
iscomprised
of
Thesystem
21.DESCRIPTION
13-21. DESCRIPTION. The system is comprised of
two sensing elements in each fuel bay (thru serial
21062273), control monitor, located inside the right
cabin wing root area, two quantity indicators located
in a cluster on the Instrument panel, and associated
wiring. Beginning serial 21062274, the dual sensing
elements have been changed to variable resistive
single element type, one in each bay, which eliminates
the control monitor. Refer to Section 16 for operation,
removal, installation, and calibration.
h. Reverse preceding steps for installation. Prior
to reinstalling equipment removed for access, service
fuel bays and check for leaks.
13-22. REMOVAL AND INSTALLATION OF SYSTEM
COMPONENTS. Refer to Section 16 for procedures.
(7) are free.. Retain them.
c. Lift off brass washer (9).
d. Mark cover (4) and body to assure later alignment of parts and remove screws (3).
e. With fine emery paper, sand off any burrs or
sharp edges on rotor shaft (21). Apply petrolatum to
rotor shaft as a lubricant, then work cover off shaft.
f. Drive hack 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 for defects.
g. Remove burrs or sharp edges on rotor shaft (21),
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.
13-23.
FUEL RESERVOIRS.
(Thru 21064535.)
13-24. DESCRIPTION. There are two reservoirs
installed in the lower fuselage, one on each side of
the aircraft, immediately outboard of the selector
valve. Each reservoir has four fuel line connections;
two from the fuel bay, one to the selector valve and
one from the selector valve, utilized for vapor return.
A drain valve is installed in the bottom of each reservolr for draining trapped water and sediment from the
fuel system.
13-25. REMOVAL AND INSTALLATION
a. Place selector valve in "OFF" position.
b. Drain all fuel from wing bay, reservoir and lines
for the reservoir being removed. (Observe precautions in paragraph 13-3.)
c. Remove front seat, carpeting and plates as necessary to gain access to reservoir.
d. Disconnect and cap or plug all fuel lines at reservolr.
e. Remove screws securing tank mounting legs to
fuselage structure.
f. Lift reservoir out.
g. Reverse the preceding steps for installation.
Prior to reinstalling equipment removed for access,
service fuel bays and check for leaks.
13-26.
FUEL SELECTOR VALVE.
(Thru 21064535.)
13-27. DESCRIPTION. A three position fuel selector valve is located In the lower fuselage between the
pilot and copilot positions. The positions on the placard are labeled "OFF, LEFT ON and RIGHT ON."
Valve repair consists of replacement of seals, springs
balls and other detail parts. Figure 13-7 illustrates
the proper relationship of parts and may be used as a
guide during disassembly and assembly.
13-28. REMOVAL AND INSTALLATION.
a. Drain all fuel from wing bays, reservoir tanks,
strainer and lines. (Observe precautions in paragraph 13-3.)
b. Remove selector valve handle.
c. Remove pedestal cover.
d. Remove access plates in floorboard and fuselage
skin in area of selector valve.
e. Disconnect and cap or plug all fuel lines at vale.
f. Disconnect square shaft from valve by removing
attached roll pin.
g. Remove bolts or screws attachg valve to sup-will
port bracket and remove valve.operate
13-14
13-29. 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 rotor
shaft (21). As yoke is lifted off, balls (8) and springs
CAUTION
Install all parts in the relative position illustrted in figure 13-6, otherwise the valve will
not operate correctly.
j. Install O-ring (18) in body rotor shaft hole. Install O-rings (19) and teflon seals (20), then slide
rotor shaft into place. Position rotor in exact relative position shown in figure 13-6, then install 0ring (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 (21) over shaft.
Remove .169" diameter pins and, readjusting rotor
(12) vs. rotor shaft (21) position as necessary, tap
roll pin (13) into place, letting it protrude on the
side illustrated.
NOTE
This roll pin (13) serves also as a stop, limiting valve rotor shaft travel.
1. Install O-ring (10) in cover (4), lubricate rotor
InstallO-ring (10) in cover(4) lubricate rotor
shaft (21) with petrolatum, install large O-ring (11)
in cover (4) and slide down into place.
CAUTION
Make sure cover (4) is installed in relative
position illustrated. A lug on the cover
serves as a stop detent and if the cover is
not
valve
installed correctly, the
not
properly.
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
13-30.
AUXILIARY FUEL PUMP.
13-31. DESCRIPTION. An electric auxiliary fuel
pump is located immediately forward of the left
fuel reservoir. An integral bypass and check valve
incorporated in the pump assembly permits fuel
flow through the pump even when 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. The auxiliary pump is used in engine
starting and in the event of engine-driven pump
malfunction.
13-32. REMOVAL AND INSTALLATION.
a. Place fuel selector valve in "OFF" position.
b. Drain fuel from pump. lines and strainer with
quick-drain control.
c. Ensure master switch and pump switch are in
"OFF" position.
d. Remove pilot's seat, carpeting and plates at
left side of pedestal as necessary for access to
pump.
e. Disconnect and cap or plug all fuel lines and
electrical connections at pump. (Observe
precautions in paragraph 13-3.)
f. Loosen the two securing clamps and lift pump
out.
g. Reverse the preceding steps for installation.
Prior to reinstalling equipment removed for access.
place selector valve to "ON" position and check for
leaks and proper pump operation.
13-33. AUXILIARY FUEL PUMP CIRCUIT. The
auxiliary fuel pump switch is a yellow and red
split-rocker type switch. The yellow right half of
the switch is labeled "START," and its upper "ON"
position, is used for normal starting and minor
vapor purging during taxi. 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. 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 at a high
capacity to supply sufficient fuel flow 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.
Maximum fuel flow is produced when the left half
of the switch is held in the spring-loaded "HI"
position. In the "HI" position, an interlock within
the switch automatically trips the right half of the
switch to the "ON" position. When the springloaded left half of the switch is released, the right
manually returned to the OFF position. When the
13-16
Revision 2
engine-driven fuel pump is functioning and the
auxiliary fuel pump is placed in the "ON" position.
a fuel/air ratio considerably richer than best power
is produced unless the mixture is leaned. If the
auxiliary fuel pump switch is accidentally placed
in the "ON" position with the master switch "ON"
and the engine stopped. the intake manifolds will
be flooded. A throttle shaft-operated microswitch
adds a resistance to the high circuit to slow down
the pump when the throttle is retarded to prevent
an excessively rich mixture. Refer to paragraph
13-34 for rigging instructions.
13-34. RIGGING THROTTLE-OPERATED MICROSWITCHES. (Refer to figure 13-7.) These aircraft
are 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 aircraft).
b. Mark position of throttle control at instrument
panel and shut down engine.
c. Remove cover (1) and adjust cam (3) to
activate fuel pump switch (6) at throttle position
marked in step "b".
d. With mixture control in "IDLE CUT-OFF "
|
electrical fuel pump switch in "ON. " and master
switch in "ON" position. listen for change in sound
of electric fuel pump as the throttle is retarded to
the marked position.
13-35. AUXILIARY ELECTRIC FUEL PUMP
FLOW RATE ADJUSTMENT. (Refer to figure 138.)
WARNING
During this test, raw fuel will drain from
the engine compartment, therefore, proper
safety precautions should be taken. Conduct
test in well ventilated area, use drip pans,
insure aircraft is properly grounded, and
keep ignition source, (cigarettes, lighters,
matches, etc.) away from area.
NOTE
These tests are to be conducted with the
supplied to the aircraft bus.
MODEL 210 & T210 SERIES SERVICE MANUAL
3
2
1. High-Boost Resistor (#1)
2. Low-Boost Resistor (#2)
Figure 13-9.
Battery Box Support
Firewall
Auxiliary Fuel Pump Resistors
a. Apply an external source of 27.75VDC ± .25V to
the aircraft bus.
b. Set mixture control at "FULL RICH."
c. Turn master switch "ON," and 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 (1) 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
respond.
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.
13-36. MAXIMUM HIGH BOOST CHECK. To
verify high position function, momentarily depress
13-18
6.
7.
3. Spacer (Typical)
4. Adjustable Slide
5. Bracket Assembly
spring-loaded rocker and verify a noticeable
increase in indicated fuel flow on the fuel flow
gage.
13-37.
FUEL STRAINER.
(Thru 21064535.)
13-38. DESCRIPTION. The fuel strainer is located
in the nose wheel well and is readily accessible
with the nose gear doors open. The strainer is
equipped with a quick-drain valve which provides
a means of draining trapped water and sediment
from the fuel system. The quick-drain control is
located adjacent to the oil dipstick.
NOTE
The fuel strainer can be disassembled,
cleaned and reassembled without
removing the assembly from the aircraft.
Beginning with T210, 21063661 thru
21064535 and those aircraft modified
by SK210-93 the fuel strainer is in-
MODEL 210 & T210 SERIES SERVICE MANUAL
sulated. The insulation material
consists of a split top and a bowl
covering. This insulation material
must be removed prior to disassembly
and reinstalled upon reassembly of the
fuel strainer.
13-39. DISASSEMBLY AND ASSEMBLY. (Refer
to figure 13-9.)
a. Place fuel selector valve in "OFF" position.
b. Open landing gear doors.
c. Drain fuel from strainer with quick-drain
control. (Observe precautions in paragraph 13-3.)
d. Disconnect strainer drain tube and remove
safety wire. nut and washer at bottom of filter bowl
and remove bowl.
e. Carefully unscrew standpipe and remove.
f. Remove filter screen and gasket. Wash filter
screen and bowl in solvent (Federal Specification
P-S-661 or equivalent) and dry with compressed
air.
g. Using a new gasket between filter screen and
top assembly, install screen and standpipe.
Tighten standpipe only finger tight.
h. 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.
i. Place selector valve in "ON" position, close
strainer drain and check for leaks and proper
operation.
j. Safety wire bottom nut to top assembly. Wire
must have right hand wrap, at least 45 degrees.
13-40. REMOVAL AND INSTALLATION,
a. Place selector valve in "OFF" position.
b. Open landing gear doors.
c. Drain fuel from strainer and lines with quickdrain control.
d. Disconnect and cap or plug all fuel lines at
strainer. (Observe precautions in paragraph 13-3.)
e. Loosen clamp and clamp bolt attaching quick. Disconnect primer line.(
g. Remove attaching bolts and remove strainer.
h. Reverse preceding steps for installation.
Place selector valve to "ON" position and check for
leaks and proper operation of quick-drain valve.
13-41. FUEL SYSTEM.
BEGINNING WITH 21064536
13-42. DESCRIPTION. The fuel system is essentially a gravity-flow system from the bay outlets to the
selector valve and a pump augmented system from
the selector valve to the engine. The fuel system is
comprised of wing bays, a selector valve, fuel
strainer, and associated plumbing. Fuel bag outlets
are located at the inboard end of the bags. A single
fuel supply line is routed down the rear doorposts to
the fuel selector valve. A fuel supply line, Interconnected with a vent line, and a separate drain line
are routed down the front doorposts. A combination
drain, and vent line is routed down the left, forward,
doorpost, from the vent crossover line to the reservoir. The fuel bays are vented by a crossover vent
line, wing tip vents, and vented fuel caps.
13-20
The upper segment of the three position (LEFT ON,
BOTH ON, RIGHT ON) fuel selector valve handles
fuel from the bays. The lower segment handles
vapor, along with returned and excess fuel from the
engine-driven fuel pump.
The reservoir accepts fuel from the selector valve,
bay drain and vent lines. The fuel flows from the
reservoir through a by-pass in the auxiliary fuel
pump (when the pump is not in operation) to the fuel
ON-OFF valve.
The fuel ON-OFF valve provides a means of stopping
fuel flow to the STRAINER and the engine driven fuel
pump. The fuel ON-OFF control is mounted on the
left side of the pedestal.
The fuel STRAINER, mounted on the firewall incorporates a remote drain valve. This valve, is mounted
on the lower, left, engine cowling. The drain val.-e
is activated by the fuel sampler cup.
13-43. FUEL SELECTOR VALVE.
13-13.)
(See figure
13-44. DESCRIPTION. A three position, six port
fuel selector valve is located beneath the floorboard.
A shaft links the fuel selector valve to a handle
mounted on the pedestal structure. The positions
of the handle are labeled "BOTH ON, LEFT ON,
RIGHT ON". Valve repair is limited to replacement
of component parts only. Figure
illustrates
the proper relationship of parts and may be used as
a guide during disassembly and assembly.
13-45. REMOVAL AND INSTALLATION.
a. Drain all fuel from wing bays, reservoir,
strainer and lines. (Observe precautions in parab. Remove selector valve handle.
c. Remove pedestal cover.
d. Remove center access plate.
d. Remove center access plate.
e. Tag, and then disconnect or plug all six lines at
valve
f. Remove screws attaching elevator cable bracket
to valve.
g. Remove nuts, washers, and bolts attaching
valve to its bracket.
h. Remove valve.
1. Reverse preceding steps for installation. Prior
to reinstalling equipment removed for access, secure
fuel bays and check all lines and fittings for leaks in
all selector valve positions.
13-46. DISASSEMBLY, REPAIR AND REASSEMBLY.
a. Remove pin (31) and shaft (30).
b. Remove spring retainer (24) spring (23) packing
(22) and seal (21) from each part of the lower body
(20)
c. Remove screw (2) holding upper body (4) and
lower body (20) together.
d. Remove lower body (20) with twisting motion.
Remove and tag washer(s) (16).
e. Cover upper body (4) and detent insert (17) with
a clean shop cloth.
MODEL 210 & T210 SERIES SERVICE MANUAL
14
12
8
11
BEGINNDIG WITH 21064536
10
1.
2.
3.
4.
5.
6.
7.
8.
Right-Hand Fuel Line
Crossvent Line
Left-Hand Fuel Line
Vent Line
Fuel Line
Drain Line
Reservoir
Auxiliary Fuel Pump
9.
10
11.
12.
13.
14.
15.
16.
ON-OFF Valve
Strainer Drain Valve
Fuel Strainer
Fuel Selector Valve
Vent Line Drain Valve
Drain Line
Fuel Line
Vent Line
Figure 13-12. Fuel System.
13-22
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
The shop cloth will contain ball (15) and
spring (14) when detent insert (17) is
removed.
f. Carefully pry detent insert (17) from upper
body (4).
g. Remove ball (15) and spring (14) from shop
cloth,
h. Remove stop pin (3) from rotor (13).
I. Cover upper body (4) completely with a clean
shop cloth,
NOTE
The shop cloth will contain seals (12),
packings (11), washers (10) and springs
(9) when the rotor is removed.
j. Push the rotor (13) out of the upper body (4).
k. Remove the rotor (13), seals (12), packings (11),
washers (10), and springs (9) from the shop cloth.
1. Check detent holes in detent insert (17) for excessive wear.
m. Replace all seals and packings.
n. Insert rotor (13), in upper body (4), place detent insert (17), over rotor (13), place washer (16)
in lower body (20), place lower body (20), over
rotor (13) insert three screws (2) and torque to
30 lbs-in. Check end play between rotor and valve
bodies.
If end play is:
(1) . 008 or greater, add S-1358-11 and/or
S-1358-12 washers to decrease end play to .001
to .007.
(2) . 007 to . 004 add (1) S-1358-12 washer.
(3) . 003 or less, disassemble valve and reassemble with different parts, recheck end play.
o. When end play is within tolerance disassemble,
retain washers.
t. Remove the upper body (4) from bench vise or
support.
u. Insert stop pin (3) Into rotor shaft.
v. Place detent insert (17) on rotor (13) with slots
for ball (15) towqrd upper body (4).
w. Place ball (15) on spring (14) align one of the
slots, with the ball (15) and depress the ball (15).
While pushing the detent insert (17) toward the upper
body (4) as the ball (15) enters the slot the detent
insert (17) may be pushed on to rotor (13) until it is
flush with the upper body (4). Rotate the detent
insert (17) until all four of its bolt holes align with
four of the holes on the upper body (4).
x. Roll packing (18) over end of rotor (13) and push
into cutout between rotor (13) and detent insert (17).
Packing (18) must not protrude beyond lip of detent
insert (17). Care must be exercised to avoid damage
to packing.
y. Place packing (19) in groove on outer edge of
detent insert (17).
z. Place lower body (20) over rotor (13). The five
bolt holes in the lower body (20) must align with the
five bolt holes in the upper body (4).
13-47. LEAK TEST
a. .With valve assembled remove stop pin (3).
b. Set valve in a closed position.
c. Apply 6-10 psi Stoddard solvent to each port
separately.
d. Maximum internal leakage 10 drops per minute.
No external leakage allowed.
13-48. ALTERNATE METHOD.
a. With valve, assembled remove stop pin (3).
b. Set valve in a closed position.
c. Apply 6-10 psi air to each port while valve is
submerged in water.
d. Maximum internal leakage equivalent to 10 drops
per minute Stoddard solvent. No external leakage
allowed.
Add two drops of Locktite 242 to end of each spring
retainer (24) after pressure test.
NOTE
13-49.
FUEL RESERVOIR
(See figure 13-14.)
Reassembly of the selector valve is facilitated by mounting upper body (4) in a bench
vise or equivalent bench support making
certain upper body (4) is protected from
damage. Fabrication of spring compressors (32) three required is necessary.
13-50. DESCRIPTION. There is one reservoir installed in the lower fuselage, on the pilot's side outboard of the fuel selector valve. The reservoir has
four fuel line connections; one from the fuel selector
valve, one from the lower right hand crossover drain
crossover drain line and
line, one from the left hand
one to the engine by way of the auxiliary fuel pump,
p. Place upper body (4) upside down in bench vise
or support.
ON-OFF valve and fuel strainer. A drain valve is
q. Replace packing (6). Lubricate spring (14) with installed in the bottom of the reservoir for draining.
petrolatum
installed in the bottom of the reservoir for draining.
(13).
and insert
insert in
in rotor
rotor (13).
petrolatum and
REMOVAL AND INSTALLATION.
r. Insert spring (9) and compress with spring compressor (32) then insert washer (10), packing (11)
13-51
in parastrainer
and lines.
lines. Observe
Observe reservoir
and seal (12). The concave portion of the seal must
in paraprecautions
strainer and
_. .13-3)
fit the convex surface of the rotor (13). Complete graph
this
b. Remove carpeting and access plate.
this for
for each
each port.
port.
at the
cap
or and plug all fuel
Remove carpeting
s. While holding the three springs (9) with the
c. Disconnect and cap or plug all fuel lines at the
and/or
(7)
washers
place
(32),
compressors
spring
spring compressors (32), place washers (7) and/or
reservoir
screws securing mounting legs to use(8) on the shaft end of rotor (13) and insert rotor (13) Remove
lage.
into the upper body (4). The seals (12) must fit flush
against the rotor (13). Release the spring compressors (32).
13-25
MODEL 210 & T210 SERIES SERVICE MANUAL
Figure 13-14.
13-26
Fuel Reservoir
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
e. Lift reservoir out.
f. Reverse the preceding steps for installation.
Prior to replacing the access plate, secure fuel bays
and check all connections for leaks.
Reassembly of valve is facilitated by mounting
in a bench vise or equivalent bench support,
making sure valve body (19) is protected from
damage. Fabrication of a spring compressor
is recommended before reassembly. Replace
packings (21) and (18) whenever rotor (17) is
removed from valve body.
NOTE
The clearance between the elevator cables
and the drain line is .37 inch minimum
and. 50 maximum.
Lower Right Hand Crossover Drain Line
From Fuel Selector Valve
Left Hand Crossvent Drain Line
To Engine
13-52.
FUEL ON-OFF VALVE.
(See figure 13-15).
13-53. DESCRIPTION. The fuel ON-OFF .alve is
a two position valve located just forward of the auxiliary fuel pump under the pilot's floorboard. The valve
control knob is located on the left lower area of the
pedestal. Valve repair consists of replacement of
component parts.
f. Ensure all component parts are clean. then coat
sparingly with lightweight oil.
g. Install new packinu ilo minto recess at top uf
valve body (19).
h. Insert spring (20) into valve body (19).
i. With spring compressor, compress spring (20).
j. Install washer (21), new packing (22). and seal
(23) into port.
k. Holding spring (20) compressed, carefully insert
rotor (17) into valve body (19), release spring compressor, and visually inspect assembly for proper
seating of seal (23) to rotor.
1. Lubricate spring (16) and ball (15) with Petrolatum.
m. Insert spring (16) into rotor (17).
n. Place ball (15) on top of spring (16).
o. Position cover (14) on valve body and turn rotor
(17) as required to index one of detents in cover.
p. Secure cover (14) to valve body (19) with screws
(13).
q. Test rotation of rotor (17) for ease of operation
and positive detent engagement.
13-54. REMOVAL AND INSTALLATION.
a. Drain all fuel from wing bays, reservoir, strainer and lines. (Observe precautions in paragraph
13-3).
b. Remove carpeting and access plate.
c. Remove control cable from clamp on valve and
control wire from valve arm.
d. Disconnect and cap or plug both the inlet and
outlet fuel lines.
~outlet
fuel
lines.~13-56.
FUEL STRAINER. (See figure 13-16.)
e. Remove bolts from bracket and remove valve.
f. Reverse the preceding steps for installation.
DESCR
ON. The fuel strainer is located
of he f
irewall. It is acceson the left f
d se
Prior to replacing the access plate, service the fuel
an hekalonetinfrlek.
honvalve
the left forward side of the firewall. It is accesbays and check allbarrows
connections
for leaks. The
a position.
check . ti o asible through the left cowl flap opening or
from
The
fuelabove
must also be checked for positive on ands off
by removing the upper engine cowling. The fuel
NOTE
strainer incorporates a quick drain valve. The valve
protrudes from the lower left side of the engine cowlWhen installing the valve make certain the
arrow on the valve points with the direction
of normal fuel flow. (Toward the engine).
13-55. DISASSEMBLY, REPAIR AND REASSEMBLY.
a. Remove screws (13) securing cover (14) to valve
body (19); carefully remove cover.
b. Remove ball (15) and spring (16) from rotor (17).
c. Slowly withdraw rotor (17) from valve body (19).
NOTE
Removal of rotor (17) from valve body (19)
will allow seal (23), packing (22) washer
(21), and spring (20) to pop free.
d. Remove seal (23), packing (22), washer-(21), and
spring (20) from valve body (19).
e. Remove packing (18) from valve body (19).
ig.
NOTE
The fuel strainer can be disassembled,
cleaned and reassembled without removing the assembly from the aircraft.
13-58.
DISASSEMBLY, ASSEMBLYAND
REASSEMBLY.
a. Place ON-OFF fuel control in OFF position
d. Drain fuel from strainer and lines with drain
valve (16).
c. Disconnect strainer drain line (10) from strainer
bowl (6) and drain valve (16).
d. Remove nut (9), step washer (8) and 0-ring (7)
at bottom of bowl (6) and remove bowl (6) remove
0-ring (5).
e. Carefully unscrew Standpipe (4) and remove.
13-27
MODEL 210 & T210 SERIES SERVICE MANUAL
3
2
33
MODEL 210 & T210 SERIES SERVICE MANUAL
8
BEGINNING WITH 21064536
1.
2.
3.
4.
Top
Gasket
Filter Screen
Standpipe
5. O-Ring
6. Bowl
7. O-Ring
8.
9.
10.
11.
10 11
12
Step Washer
Nut
Drain Line
Nut
13
12. Washer
13.
14.
15.
16.
17.
18.
Bracket
Fitting
O-Ring
Drain Valve
Washer
Screw
14
Figure 13-16.
Fuel Strainer.
13-29
MODEL 210 & T210 SERIES SERVICE MANUAL
f. Remove filter screen (3) and gasket (2). Wash
filter screen and bowl in solvent (p-S-661) and dry
with compressed air.
g. Using a new gasket (2) install filter screen (3)
and standpipe (4). Tighten standpipe finger tight.
h. Using new O-rings (5) and (7) install bowl (6).
The step washer (8) must be installed so that the step
seats against the O-ring (7), connect drain line (10).
i. Place ON-OFF fuel control in ON position.
j. Check for fuel leaks.
k. Check drain valve (16) for operation.
13-59.
VENTED FUEL FILLER CAPS.
13-60.. DESCRIPTION. The filler cap assemblies
may be constructed of either metal or red plastic.
Both cap assemblies incorporate a vent safety valve
that provides vacuum and positive pressure relief
for their respective fuel tanks. It is important that
both type caps to be cleaned on as required basis,
if proper filler cap sealing is to be maintained.
13-61. METAL "FLUSH-TYPE" FILLER CAPS.
Except for minor differences in construction and
weight, metal fuel filler caps perform the same
function as red plastic fuel filler caps. The caps
are interchangeable and will fit the same adapter
assembly.
13-62. INSPECTION.
NOTE
If fuel collects in the handle well it could indicate stem O-ring leakage. Fuel collecting
around perimeter of cap could indicate cap
O-ring or check valve leakage.
a. Remove fuel cap from adapter (7), remove safety
chain (9) from cap and cover or plug fuel opening to
keep out foreign matter.
b. Remove nut (10) and, oberving position of lock
plate (6) in relation to stem (14) disassemble cap.
c. Note resiliency of 0-rings (3 & 13) and condition of grooves. If the 0-rlngs (3 & 13) have deteriorated they must be replaced
13-63. CLEANING.
a. Using a cotton swab and Stoddard solvent or
equivalent, gently lift edges of rubber umbrella (5)
and clean stainless steel seat and umbrella removing
all contaminates. Using a second swab wipe seat
and umbrella thoroughly, removing all cotton fibers.
Repeat until swabs show no discoloration.
b. If O-ring grooves appear contaminated, clean
with Stoddard solvent or equivalent and cotton swabs.
c. Ascertain that all vent holes in check valve are
unobstructed.
d. Clean cap body and lock plate, check for defects.
e. If the umbrella continues to leak or is deteriorated it must be replaced.
f. To remove umbrella, lubricate the umbrella
tearing the stem.
13-30
g. To replace the umbrella, lubricate the umbrella
stem with (MIL-H-5606) hydraulic fluid and use a
small blunt tool to insert the retaining knob on the
umbrella stem into the check valve body to prevent
damaging the stem.
13-64. REASSEMBLY.
a. Place split washer (16) in cap well correctly.
b. With handle (1) and O-ring installed on stem (14),
insert stem (14) through split washer (16) on cap
body (2).
c. Place spring (15) on stem (14).
d. Position cap handle (1) to full "OPEN" position.
e. Place lock plate (6) on threaded end of stem (14)
and align all three lugs (12) with three guide bosses
on the cap body (2).
f. Check that square hole in bottom of lock plate
(6) is aligned with square surface on threaded end of
stem (14).
NOTE
It is possible to install the lock plate (6)
180º out of the desired position, if the alignment procedures in steps "d" and "Y" are
not followed. If the cap will not fit when assembled, remove the lock plate (6) and rerotating it 180".
adapter
assemble
g. Compress the lock plate (6) and fuel cap body (2)
and secure with washer (11) and nut (10).
h. Connect fuel cap assembly to safety chain (9) and
reinstall in tank.
13-65. RED PLASTIC "FLUSH-TYPE" FILLER
CAPS. A red plastic "Flush-Type" vented filler cap
may be used. Extra care is required when reinstalling plastic filler caps in the fuel filler adapter assembly. An improperly installed filler cap could cause
a loss of fuel from the tanks during flight.
13-66. INSPECTION.
NOTE
If fuel collects in the handle well it could indicate stem O-ring leakage. Fuel collecting
around perimeter of cap could indicate cap
outer seal or check valve leakage.
a. Remove fuel cap from adapter (8), remove safety chain (10) from cap and cover or plug fuel opening
to keep out foreign matter.
b. Rotate cap handle (1) to the "OPEN" position,
compress cap body (2) and lock plate (6) to expose
the . 125 inch diameter handle pin (17).
c. Using a small wire push out the handle pin (17).
d. Note resilience of O-ring (13) and outer seal (3)
and condition of grooves. If the O-ring (13) or the
outer seal (3) have deteriorated they must be replaced.
e. Note condition of tabs on lock plate (6) for signs
of abnormal wear, if such wear is evident replace
the complete cap assembly.
MODEL 210 & T210 SERIES SERVICE MANUAL
1.
2.
3.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
Handle
Fuel Cap Body
O-RingCheck
Valve (Vent)
Umbrella
Fuel Cap Lock Plate
Adapter Assembly
Placard
Safety Chain
Nut
Washer
Lug
O-Ring
Stem
Spring
Split Washer
Handle Pin
* A Letter M on the
fuel cap body located
under the handle (1),
signifies that the 0-ring
(3) mounting groove
is machined.
Vent safety valve (4) opens at or before
.25 PSI vacuum, and 5.0 PSI pressure.
16
3
4
15
14
12
Figure 13-17. Fuel Filler Cap-Metal (Sheet 1 of 2)
Revision 2
13-31
MODEL 210 & T210 SERIES SERVICE MANUAL
13-67. CLEANING.
13-69. LEAK TESTING METAL OR RED PLASTIC
a. Using a cotton swab and Stoddard solvent or
FILLER CAPS. The following procedure may be
equivalent, gently lift edges of rubber umbrella (5)
used to detect fuel filler cap leakage.
and clean stainless steel seat and umbrella removing
a. Service the aircraft with approved fuel, filling
all contaminates. Using a second swab wipe seat and
each fuel bay.
umbrella thoroughly, removing all cotton fibers.
b. Place the fuel selector in the OFF position.
Repeat until swabs show no discoloration.
c. Plug one of the fuel bay vent lines (where it prob. If 0-ring or outer seal grooves appear contamitrudes beneath the wing) with a small rubber plug or
nated, clean with Stoddard solvent or equivalent and
tape.
cotton swabs.
d. Connect a rubber hose to the other vent. Then
c. Ascertain that all vent holes in check valve are
tee into this hose a pressure measuring device, such
unobstructed.
as a water manometer, manifold pressure gage or
d. Clean cap body and lock plate, check for defects.
airspeed indicator.
e. If the umbrella continues to leak or is deterioe. Blow into the open end of the hose. The pressure
rated it must be replaced.
must not exceed .7 psi which equals 20 inches of
f. To remove umbrella, lubricate the umbrella
water on a water manometer. or 1. 43 inches Hg on a
stem with (MIL-H-5606) hydraulic fluid to prevent
manifold pressure gage, or 174 kts on an airspeed
tearing the stem.
indicator.
g. To replace umbrella, lubricate the umbrella
WARNING
item with (MIL-H-5606) hydraulic fluid and use 2
small blunt tool to insert the retaining knob on the Do not inhale fuel vapor while blowing into
umbrella stem into the check valve body to prevent
the
rubber hose.
damaging the stem.the
rubber hose.
13-68.
f. It may take several applications of pressure to
bring the bay to the desired pressure.
REASSEMBLY.
WARNING
NOTE
If fuel was observed leaking around the cap
periphery prior to disassembly and the leakage was not due to a bad O-ring or outer seal
an additional split washer (16) may be added
for a total of two, prior to reassemblying
cap. To make sure that these washers are
not installed upside down, check to see that
edges of the split parallel the respective
sides of the cap well The addition of a
washer under the cap handle will increase
the effort required to uncap the fuel tank.
b. Install fuel cap body (2) on stem (14).
c. Check that three metal plates (12) on top rim of
lock plate (6) are aligned with three guide bosses on
fuel cap body (2).
Do not apply regulated or unregulated air
pressure from an air compressor to the
fuel vent. Over inflation and major structural damage will occur if more than .7 psi
is applied.
g. Pinch or close the rubber hose to sustain pressure in the fuel bay.
h. Apply a soap solution to the fuel filler caps and
inspect for leakage around the rubber seal to filler
neck junction, the fuel cap vent, and the fuel cap
handle stem. Load the cap sideways in all directions
by pressing on the fuel cap vent housing by hand.
NOTE
No leakage is permissible.
CAUTION
or repair in accordance with Cessna
Service Information Letter SE80-59,
Supplement #1, dated, June 23, 1980.
It is possible to install the handle pin in the
pin hole 180 ° out of the desired position, if
the alignment procedure in step "c" is not
followed. If the handle (1) is not installed^^
properly the FWD arrow on the cap will not
align with the arrow on the placard (9) when
the cap is reinstalled.
d. Compress cap body (2) and lock plate (6), install
split washers cap body (2) and lock plate (6) install
e. Install cap handle (1) on stem (14) so that the
handle (1) will be in the open position.
f. Insert
handle
(1) will
inthe
position.
(tape
f. Insert handle pin (17) open
through handle (1) and
stem (14).
g. Connect fuel cap assembly to safety chain (10)
and reinstall fuel cap. Make certain that the arrow
on the fuel cap body (2) and the arrow on the placard
(9) align.
If leaks are
CAUTION
Care must be exercised in removing the
fuel filler caps until the system has been
depressurized.
i.
After replacement of either fuel filler cap. repeat
the Inspection.
j. Remove the rubber hose, unplug or remove the
from the other fuel vent, and place the fuel selector in the desired position.
13-33/(13-34 blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 14
PROPELLER AND GOVERNOR
WARNING
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the
propeller, since a loose or broken wire or a
component malfunction could cause the propeller
to rotate.
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
2H6/14-1
PROPELLER ............
2H8/14-1
Description ...........
.2H6/14-1
Repair ............
.. . .2H77/14-2
Trouble Shooting ..
2H9/14-2B
Removal ...........
.2H9/14-2B
Installation ..........
Time Between Overhaul (TBO) . . 2H10114-3
.2H10/14-3
.......
GOVERNOR ....
Description ..........
.2H10I4-3
14-1.
PROPELLER.
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 governpr. An increase or decrease in throttle
setting or a change in aircraft attitude will affect the
balance which maintains a given RPM. If the throttle
is opened further or if aircraft speed is increased,
engine RPM will try to increase. The governor
senses this and directs oil pressure to the forward
side of the piston. The blades will be moved to a
2H13/146
Trouble Shooting .......
...... 2H13146
Removal
Control Arm and Bearing Assembly. 2H13/14-6
2H13/14-6
.....
Removal and Installation.
2H13/14-6
Governor Installation .......
...... 2H14/14-7
High-RPM Adjustment
2H1414-7
Rigging Governor Control .....
Time Between Overhaul (TBO) * · 2H15/14-8
higher pitch and engine speed will remain constant.
Conversely, if the throttle opening or the aircraft
speed is decreased, the engine RPM will try to decrease. The governor senses this and allows oil to
drain from the forward side of the piston. Spring
tension and centrifugal twisting moment will move
the blades to a lower pitch to maintain the selected
engine speed.
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. 31rerations and who may accomplish them. When .aicing repairs or alterations to a propeller FAR 43.
FAA AC No. 43.13 and the propeller manufacturer's
instructions must be observed.
Revision 3
14-
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
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.
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 para 14-6, e.
Defective seals, gaskets,
threads, etc., or incorrect
assembly.
Propeller repair or replacement
is required.
FAILURE TO CHANGE PITCH
FULLY.
ENGINE SPEED WILL NOT
STABILIZE.
OIL LEAKAGE AT PROPELLER MOUNTING FLANGE.
OIL LEAKAGE AT ANY
OTHER PLACE.
14-2
Connect or re-
MODEL 210 & T210 SERIES SERVICE MANUAL
THIS PAGE INTENTIONALLY LEFT BLANK
14-2A blank
14-2A blank
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
WARNING
^Avoid
WARNING
Be certain that magneto is GROUNDED
before turning propeller.
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), which attach the
spinner bulkhead to the propeller.
14-6.
INSTALLATION.
a. If the spinner bulkhead was removed, position
bulkhead so the propeller blades will protrude thru
the spinner with ample clearance. Install spinner
bulkhead attaching screws (10), which attach the
spinner to bulkhead.
CAUTION
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.
NOTE
NOTE
If aircraft is configured with optional propeller anti-ice system, the slip ring assembly must be installed with or prior to propeller. Use care to prevent damaging
brushes and slip ring, and insure proper
alignment. Reconnect slip ring wires according to applicable wiring diagram.
NOTE
WARNING
If the optional propeller anti-ice system
is installed, use caution when removing
propeller. Removing the propeller without
the anti-ice slip ring requires disconnecting
nine wires at spinner bulkhead, since the
slip ring is mounted to the bulkhead. Wires
should be identified according to wiring diagram to facilitate reassembly. During removal. installation, or other maintenance,
use care to prevent damaging slip ring and
brushes.
14-2B
Be certain that magneto is GROUNDED
before turning propeller.
*
MODEL 210 & T210 SERIES SERVICE MANUAL
*~ *d.
Lubricate the hub mounting studs with A-1637-16
(MIL-T-83483) grease.
CAUTION
ALL PROPELLER STUDS AND NUTS ARE
REQUIRED TO BE INSTALLED WITH
LUBRICATION ON THE HUB MOUNTING
STUDS.
e. 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.
f. Install propeller attaching washers and new nuts (9)
and work propeller aft asfar as possible, then tighten
nuts evenly.
WARNING
g.
DO NOT USE ALL STEEL LOCKNUTS. USE
ONLY NEW ELASTIC ELEMENT LOCKNUTS
WHEN INSTALLING PROPELLER.
Torque nuts 45 to 50 lb-ft. LUBRICATED TORQUE
ONLY. Refer to McCauley Service Bulletin 227, or
latest revision, as applicable for propeller stud and
. CAUTION
USE OF CROW FOOT OPEN-ENDED TORQUE
WRENCHES CAN CAUSE SLIPPAGE AND
LEAVE MARKS ON THE ENGINE OUTPUT
FLANGE IF CARE IS NOT USED DURING THE
TORQUE PROCESS.
USE PROPER CALCULATIONS WHEN USING
TORQUE ADAPTERS TO ENSURE CORRECT
INSTALLATION TORQUE.
TO PRODUCE CONSISTENT AND ACCURATE
MCCAULEY
TORQUE,
INSTALLATION
RECOMMENDS AN ADJUSTABLE "CLICK"
TYPE WRENCH WITH NON RACHETING,
INTERCHANGEABLE, 12 POINT BOX-END
WRENCH HEADS.
IT MAY BE NECESSARY TO USE VARIOUS
ADAPTERS IN CERTAIN APPLICATIONS.
IS
STRONGLY
HOWEVER,
IT
RECOMMENDED THAT EXTREME CAUTION
BE EXERCISED TO ENSURE
THAT
ACCURATE TORQUE IS BEING APPLIED
FOR MAXIMUM RETENTION.
h. Install spacers (4) and spinner support (3) on
propeller cylinder (5). If spacers (4) are not centered
mechanically (piloted), visually center and hold them
until spinner support (3) isforced firmly in place.
i. Hold spinner (1) snug against spinner support (3) and
check alignment of holes in spinner (1) with holes in
spinner bulkhead( 1). Add or remove spacers (4)
from propeller cylinder (5) until holes are within .050
of alignment.
j. Push hard on spinner (1) to align holes and install
screws and washers (if required) in three (3) or more
equal spacers around the spinner bulkhead (11).
Relax pressure on spinner and install remaining
screws and washers (if required) in spinner.
k. Tighten all screws uniformly around the spinner.
(TBO).
OVERHAUL
BETWEEN
14-6A. TIME
Propeller overhaul shall coincide with engine
overhaul, but shall not exceed limits specified in
McCauley Service Bulletin 137 and all revisions
and supplements thereto. Refer to Sections 12
and 12A for engine overhaul periods.
14-7. GOVERNOR.
14-8. DESCRIPTION. The propeller governor is a singleacting, 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 singleacting 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,
flyweight and speeder spring act together to open
and close governor oil passages as required to
maintain a constant engine speed.
NOTE
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
sentto
ON MOST AIRPLANES, A TORQUE WRENCH
CANNOT BE FITTED DIRECTLY ON THE
PROPELLER MOUNTING NUT BECAUSE OF
THE LACK OF CLEARANCE BETWEEN THE
FLANGE AND ENGINE CASE. AN ADAPTER
MUST BE USED ON THE TORQUE WRENCH.
THE USE OF A TORQUE WRENCH WITH ANY
EXTENSION REQUIRES THE
FORM OF
TORQUE READING ON THE WRENCH TO BE
CHANGED TO OBTAIN THE CORRECT
TORQUE APPLIED AT THE NUT. TO OBTAIN
CORRECT RESULTS REFER TO THE
FORMULA IN SECTION 1.
Temporary Revision Number 4
14-3
MODEL 210 &T210 SERIES SERVICE MANUAL
8
7
6
1
Additional
1. Propeller Spinner
2. Screw
3. Spinner Support
4. Spacer
5. Cylinder
6. Screw
7. Stud
8. Washer
9. Nut
10.
11.
12.
13.
14.
15.
16.
17.
Screw
Spinner Bulkhead
Engine Crankshaft
O-Ring
Dowel Pin
Propeller
Tube
Ring
Figure 14-1. Propeller installation (Sheet 1 of 2)
14-4
Revision 2
spacers
(4)
may
be
required when installing a new
spinner (1) to ensure a snug fit
between spinner (1) and support
THRU SERIAL 21062003:
(3).
Part
Number
Order
Cessna
more
NOT
USE
0752620-2. DO
than 6 spacers in this installation
SERIAL
WITH
BEGINNING
Order
21062004:
NUMBER
Cessna Part Number 0752620-3.
DO NOT USE more than 14
spacers in this installation.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
The result of rigging is full travel of the
governor arm (bottomed out against both
high and low pitch stops) with some cushion
at each end of control travel.
SHOP NOTES:
14-8
Revision 3
14-16. TIME BETWEEN OVERHAUL. (TBO)
Propeller governor overhaul shall coincide with engine
overhaul. Refer to section 12 or 12A for engine time
between overhaul (TBO) intervals. The governor and
propeller overhaul manuals are available from Cessna
Parts Distribution (CPD 2) through Cessna Service
Stations.
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 15
UTILITY SYSTEMS
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
UTILITY SYSTEMS
..........
2H18/15-2A
Heating System. ..........
.2H18/15-2A
Description ..........
.2H18/15-2A
Operation ...........
.2H18/15-2A
Trouble Shooting .......
.2H18/15-2A
|
Removal and Installation of
.Functional
Components
.......
. 2H18/15-2A
Defrosting System .........
.2H18/15-2A
Description ............
2H18/15-2A
Operation ......
..
.2H18/15-2A
Trouble Shooting ........
.2H18/15-2A
Removal and Installation of
Components
.........
.2H18/15-2A
Ventilating System .........
.2H18/15-2A
Description .........
. .2H18/15-2A
Operation .
............
218/15-2A
Trouble Shooting .........
2H19/15-3
Removal and Installation of
Components ....
.....
2H19/15-3
De-Ice and Anti-Ice Systems ..
. .. 213/15-11
Wing and Horizontal Stabilizer
One-Cycle De-ice System
(Thru 21062968) ...
. .213/15-11
Description .........
213/15-11
System Operation
....
. 213/15-11
Removal and Installation of
Components ........
213/15-11
Trouble Shooting .......
213/15-11
Operational Check ......
215/15-13
Adhesion Test ........
216/15-14
Cleaning
.
......... 216/15-14
|
Boot Protective Products . . .. 216/15-14
Approved Repairs (Cold Patch). 217/15-15
Approved Repairs (Damage to
Tube Area) ........
217/15-15
Approved Repairs (Damage to
Fillet Area) .......
217/15-15
Approved Rep-irs (Damaged
Veneer .........
.218/15-16
Materials Required for
Installation of Boots ...
218/15-16
Replacement of Boots
.218/15-16
..
Wing and Horizontal Stabilizer
Three-Cycle De-Ice System
(Beginning with 21062969) . .
. 219/15-17
Description ........
.19/15-17
System Operation. .....
.219/15-17
Flight into Known Icing Equipment
and Systems (Beginning with
21063253) .........
.
219/15-17
Description ........
.219/15-17
Wing, Horizontal Stabilizer
and Vertical Fin De-Ice
System (Beginning with
21063253) ....
.219/15-17
Description
.
..... 2114/15-22
Trouble Shooting ..........
De-Ice Flow Valve .........
Description ...........
De-Ice Flow Valve
Overhaul .............
Check
(Known Icing) ...........
BootRepair(ColdPatch) ...
De-Ice Boot Types of Damage
and Repair ..............
Materials Required for
Installation .............
Boot Replacement .........
Timer ....................
Description ............
Functional Test (Thru
1982 Models) .........
Functional Test
(Beginningwith
1983 Models) .........
Propeller Anti-Ice Boots (Known
Icing) ............
Windshield Anti-Ice Panel (Known
Icing) .
.
.......
.
Pitot Tube and Stall Warning
Heaters (Known Icing) ..
....
Description ........
Removal and Installation ..
Ice Detector Light ........
Description
.........
95-Amp Alternator Installation
Dual 60-Amp Alternator
Installation ..........
Control Surface
Dischargers ..........
Description .
.......
Resistance Check
.....
Propeller Anti-Ice System.
..
TroubleShooting .............
Slip Ring Removal ...........
Slip Ring Installation .........
Slip Ring Alignment Check ...
Removal of Propeller
Anti-Ice Timer .............
Installation of Propeller
Anti-Ice Timer .............
PropellerAnti-Ice Ammeter ...
Description ...............
Removal ..................
Installation ...............
TroubleShooting ..........
TimerTest
..................
Installation and Alignment of
Brush Block Assembly ......
ReplacementofDe-IceBoots ..
2114/15-22
2114/15-22
2114/15-22
2114/15-22
2123/15-31
2124/15-32
2124/15-32
2124/15-32
2124/15-32
2124/15.32
2124/15-32
2124/15-32
2124/15-32
. 2J1/15-33
2J1/15-33
2J1/15-33
2J1/15-33
. 2J1/15-33
2J1/15-33
2J1/15-33
. 2J2/15-34
2J2/15-34
2J2/15-34
2J2/15-34
2J2/15-34
. 2J3/15-34A
2J3/i5-34A
2J4/15-35
2J415-35
2J4/15-35
2J7/15-38
2J8/15-39
2J8/15-39
2J8/15-39
2J8/15-39
2J8/15-39
28/15-39
2J8/15-39
2J9/15-40
2J9/15-40
Revision 3
15-1
MODEL 210 & T210 SERIES SERVICE MANUAL
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
Windshield Anti-Ice Panel
(Removable) ...................
Description ..................
Removal and Installation .....
Windshield Anti-Ice Panel (Fixed)
Description ..................
Removal and Installation .....
Trapped Moisture ............
Oxygen System ..................
Description ...............
Maintenance Precautions .....
Replacement of Components ..
Oxygen Cylinder General
Information .............
15-2
Revision 3
2J10/15-40A
2J10/15-40A
2J10/15-40A
2J10/15.40A
2J10/1540A
2J10/15.40A
2J14/15-40E
2J15/15.40F
2J15/15-40F
2J16/15-41
2J16/15-41
2J18/15.43
Service Requirements ...
Inspection
Requirements .........
System Components Service
Requirements ...........
Inspection Requirements
Masks and Hose ...........
Maintenance and
Cleaning .............
System Purging ............
Functional Testing
System Leak Test .........
System Charging ..........
2J22/15-47
2J22/15-47
2J22/15-47
2J22/15.47
2J23/15-48
2J23/15.48
2J23/15-48
2J24/15-49
2J24/15-49
MODEL 210 & T210 SERIES SERVICE MANUAL
15-1.
UTILITY SYSTEMS.
15-2. HEATING SYSTEM.
(See Figure 15-1.)
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 of the
firewall, a duct across the aft side of the firewall,
a push-pull control on 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 axtending 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 aircraftwith
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
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
completely
when the controlis pushed full in
15-5. TROUBLE SHOOTING. Most of the opertional 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.
15-6. REMOVAL AND INSTALLATION OF COMPONENTS. Figures 15-1 and 15-2 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. DEFROSTING SYSTEM.
(See figure 15-1.)
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 control.
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 paragraph 15-5 for trouble shooting the
heating and defrosting system.
15-11. REMOVAL AND INSTALLATION OF COMPONENTS. Figure 15-1 and 15-2 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 defective 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.
(See figure 15-3.)
15-13. DESCRIPTION. The system is comprised of
an airscoop, mounted in the inboard leading edge of
each wing, outlet control valves, installed in overhead
consoles, located on the aircraft centerline, control
valves, located above each rear doorpost, 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 duc ting connecting the systems. On 1977 thru 1980
models, fixed inlet scoops are installed in the
lower forward cabin. The scoops are ducted to
the avionics equipment to aid in cooling, and under the cabin floor to help prevent exhaust fumes
from entering the cabin.
15-14. VENTILATING SYSTEM OPERATION. Air
received from scoops mounted in the inboard leading
edges of the wings is ducted to individually-controlled
control valves, two of which are mounted in each of
Revision 1
15-2A/(15-2B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
two overhead consoles and one mounted in a console
located above each rear door post. Each control
valve meters the incoming cabin ventilation air, and
provides an expansion chamber which reduces inlet
air noise. Filters at the air inlets are primarily
noise reduction filters. Air volume from the louvers in the outlet control valves is controlled by
knobs located on the end of each valve. Beginning
with 1982 models. outside air is routed from the wingmounted scoopsthrough valves in each wing root to
our lever-adjusted ventilators in the cabin. The leveradjusted ventilators replace the outlet control valves
and are located in the same area. Airflow from the
wing root valves is controlled by a lever in the overhead
console labeled: OVERHEAD AIR VENTS ON OFF.
Beginning with 1983 models without air conditioning,
the fresh air scoops and wing root valves are replaced
by ducts stopped withadjustable doors located on the
underside of each wing near the root. The adjustable
doors in the ducts are controlled by the lever labeled:
OVERHEAD AIR VENTS. Cabin ventilation is provided
by two fresh air scoop doors, one on each side of the
fuselage, Just forward of the front seats. The left scoop
door is operated by a knob on the instrument panel
labeled: CABIN AIR, and the right scoop door is controlled by a knob adjacent to the CABIN AIR knob
labeled: AUX CABIN AIR. Fresh air from the scoops
is routed to a duct running across the aft side of the
firewall, where it is distributed to the cabin. As long
as the CABIN HEAT knob is pushed full in, no heated
air can enter the firewall duct, however, as the CABIN
HEAT knob is gradually pulled out, more and more
heated air will blend with fresh air from the scoops.
Any ofthe knobs may be set to any desired position
to provide comfortable cabin temperatures.
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 inner tube in the control
valve could also bind or stick. requring repair or replacement of the control valves. Check the filter elements in the airscoops in the leading edges of the
wings for obstructions. The elements -nay be removed
and cleaned or replaced. Since air passing through
the filters is omitted 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. could cause a sligh increase in noise
level.
15-16. REMOVAL AND INSTALLATION OF COMPONENTS. Figure 15-3 may be used as a guide for
removal and .astallation 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 control valve should be repaired or replaced.
Check for proper operation of ventilating system controts after installation or repair.
MODEL 210 & T210 SERIES SERVICE MANUAL
.
3. Clamp Bolt
6.
7.
10
11
12
Cabin Heat Control
Defroster Nozzle
9.
10.
Cabin
ControlHeat
ArmValve
11.
13.
Spring
Valve Seat
14.
Valve Body
14
Thru 21064135
Figure 15-1.
15-4
Model 210 Heating and Defrosting System (Sheet 1 of 2)
MODEL 210 & T210 SERIES SERVICE MANUAL
7
Detail B
Beginning with 21064536
B
12
9
o
ro
mAl rtnC.9elzzoNretsorfeD.1
10.
spring
Nut2.
3.
4.
5.
6.
7.
8.
Clamp Bolt
Shaft
Valve
Cabin Heat Control
Defroster Control
Duct
11.
12.
13.
14.
15.
16.
Cabin Heat Valve
Valve Seat
Valve Body
Adapter
Tube Assembly
Shroud
Detail
Beginning with 21064136
B
21064136 thru 21064535
Figure 15-2. Model T210 Heating and Defrosting System (Sheet 2 of 2)
15-7
MODEL 210 & T210 SERIES SERVICE MANUAL
10
2
2
Detail
Detail B
B
Detail A
5
B
NOTE
Filter elements (12) are
installed in the leading
edge inlets and at the
pilot's and copilot's overhead console air valve
duct connections.
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
7.
18.
19.
20.
21.
22.
23.
24.
Clamp
Connector
Mounting Bracket
Inner Tube
Nut
Outer Tube
Wheel
Retainer
Mounting Bracket
Housing
Fresh Air Scoop
Filter Element
Tie
Overhead Console
Fitting
Hose
Cabin Air Control
Aux Cabin Air Control
Cold Air Inlet
Air Scoop Assembly
Inlet Screen
Air Valve Assembly
Scoop Door
Fuselage Skin
12
16
19
20
19
20
22
23
24
* Electronics Cooling
Thru 21064135
Figure 15-3.
15-8
13
Detail
Ventilating System (Sheet 1 of 3)
C
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
15-17.
DE-ICE AND ANTI-ICE SYSTEMS.
15-17A. WING AND HORIZONTAL STABILIZER
ONE-CYCLE DE-ICE SYSTEM. (Thru 21062968.)
15-18. DESCRIPTION. The de-ice system consists
of an engine-driven pneumatic pump, an annunciator
light to monitor system operation, a timer, control
valves, pneumatic de-icing boots, installed on the
leading edges of the wings and horizontal stabilizer
and the necessary hardware to complete the system.
CAUTIO N
Always allow sufficient ice build-up for efficient ice removal before actuating the de-ice
system. If de-ice system is actuated continuously, or before ice has reached sufficient
thickness, the ice will build up over the boots
instead of cracking off.
15-19. SYSTEM OPERATION. The boots expand
and contract, using pressure or vacuum from the
engine-driven vacuum pump. Normally, vacuum is
applied to all boots to hold them against the leading
edge surfaces. When a de-icing cycle is initiated,
the vacuum is removed, and a pressure is applied to
"blow up" the boots. The resulting change in contour
of the boot will break the ice accumulated on the
leading edges. The ice will then be removed by normal in-night air forces. Controls for the de-icing
system consist of a spring-loaded on-off rocker
switch on the left switch and control panel, a pressure indicator light on the upper left side of the instrument panel, and a 5-amp circuit breaker switch
on the left sidewall circuit breaker panel. The twoposition de-ice switch, labeled DE-ICE PRESS, is
spring-loaded to the normal off (lower) position.
When pushed to the ON (upper) position and released,
the system timer (located on the glove box) is energized which in turn activates one de-icing cycle.
Each time a cycle is desired, the switch must be
pushed to the ON position and released. The pressure
indicator light, labeled DE-ICE PRESSURE, should
come on within four seconds after the cycle is initiated and remain on for two or three seconds if the
system is operating properly.
15-20. REMOVAL AND INSTALLATION OF DE-ICE
SYSTEM COMPONENTS. For removal and installation of de-ice system components, see figure 15-4.
See figure 15-5 for ice detector light installation.
15-21. TROUBLE SHOOTING - WING AND HORIZONTAL STABILIZER ONE-CYCLE DE-ICE SYSTEM.
TROUBLE
DE-ICE BOOTS DO NOT
INFLATE OR INFLATE
SLOWLY.
DE-ICE BOOTS DO NOT
DEFLATE OR DEFLATE
SLOWLY.
PROBABLE CAUSE
REMEDY
Loose or faulty wiring.
Repair or replace wiring.
Loose or damaged hose.
Tighten or replace hose.
Loose or missing gasket.
Tighten fitting and/or replace gasket.
Shuttle valve malfunction.
Replace shuttle valve.
Pressure relief valve set too low.
Reset or replace valve.
Pressure relief valve malfunction.
Replace pressure relief valve.
Defective timer.
Replace timer.
Pressure relief valve malfunction.
Replace pressure relief valve.
Shuttle valve malfunction.
Replace shuttle valve.
Defective timer.
Replace timer.
Revision 1
15-11
MODEL 210 & T210 SERIES SERVICE MANUAL
CAUTION
The negative ground must be applied to the
black wire; red is positive. A reverse
voltage will ruin timer diode. The 28 VDC
must be filtered if it is rectified from AC.
If possible use a battery.
CAUTION
Use only the following instructions when cleaning de-ice/anti-ice boots. Disregard instructions which recommend petroleum base liquids
(MEK, non-leaded gasoline, etc.) which can
harm the boot material
15-22A. ADHESION TEST.
a Clean boots with mild soap and water, then rinse
a. Using excess material trimmed from ends of an
thoroughly with clean water.
wing or empennage de-ice boot, prepare one test
NOTE
specimen for each de-ice boot installed.
b. This specimen should be one-inch wide and four
or more inches long.
Isopropyl alcohol can be used to remove
c. Cement specimen to installation surface adjacent
grime which cannot be removed using
to installed de-ice boot, following the identical procesoap. If isopropyl alcohol is used for
cleaning, wash area with mild soap and
dure used for boot installation.
water, then rinse thoroughly with clean
d. Leave one-inch of the strip uncemented to attach
a clamp.
water.
e. Four hours or more after de-ice boot installaDE-ICE AND ANTI-ICE BOOT PROTECTIVE
tion, attach a spring scale to uncemented end of each 15-23A.
PRODUCTS. Two rubber treatment products, Age
strip and measure force required to remove the strip
Master #1, and Icex are approved for use on de-ice
at a rate of one-inch per minute. The pull shall be
boots and anti-ice boots of Cessna aircraft. Age
applied 180º to the surface. (Strip doubled back on
Master #1 protects the rubber against deterioration
itself).
from ozone, sunlight weathering, oxidation and poluf. A minimum of five pounds tension (pull) shall be
tion. Icex helps retard ice adhesion and keeps the
required to remove test strip.
boots looking new longer; both products are produced
and recommended by B. F. Goodrich. Age Master #1
NOTE
(part #74-451-127) and Icex (part #ICEX) are available from the Cessna Supply Division.
If less than five pounds is required accepta. Mask surrounding areas before applying Age
ability of the de-ice boot adhesion shall be
Master #1 to clean, dry boot surfaces. Apply with
based on carefully lifting one corner of the
a cheesecloth swab. DO NOT SPRAY this product;
de-ice boot in question sufficiently to attach
a rubbing or brushing action is required for the proa spring clamp and attaching a spring scale
tective agent to penetrate the rubber surfaces. Apply
to this clamp. Pull with force 180° to the
three or more coats allowing a 5 to 10 minute drying
surface, and in such a direction that the deperiod between applications. However, the total
ice boot tends to be removed on the diagonal
amount applied should not exceed 0. 3 to 0. 4 ounce
If a force of five pounds per inch of width
per square foot of boot surface.
can be exerted under these conditions, the
b. Mask surrounding areas before applying a light
installation shall be considered satisfactory.
coat of Icex with a cheesecloth swab to clean, dry
Width increases as corner peels back.
boot surfaces. A heavy coat of Icex will result in a
sticky surface which collects dust and dirt. One
g. Re-cement corner following installation procequart of Icex will cover approximately 500 square
dure.
feet. If boots have been treated with Age Master #1,
allow it to dry for a minimum of 24 hours before
applying the Icex. Apply Icex Spanwise in a single
continuous back and forth motion.
Failure to achieve five pounds adhesion per
inch of width requires reinstallation of the
|CAUTION
CAUTION
de-ice boot.
NOTE
Possible reasons for failure are: dirty surfaces, cement not mixed thoroughly. Corrosion of metal skin may occur if good adhesion
is not attained, especially around rivet heads
and metal skin splices. If these adhesion
requirements are met, the aircraft may be
flown immediately. Do not inflate de-ice
boots within 48 hours of installation.
15-23.
15-14
CLEANING.
Revision 1
Protect adjacent areas, clothing, and wear
plastic or rubber gloves during application.
Age Master stains clothing and Icex contains silicone which makes paint touch-up
nearly impossible. Waterless hand cleaner
is beneficial for cleaning hands, equipment
and clothing.
Age Master #1 and Icex coatings last approximately
150 hours on wing and stabilizer boots and 15 hours
on propeller boots.
MODEL 210 & T210 SERIES SERVICE MANUAL
15-24. APPROVED REPAIRS.
Scuff or Surface Damage.)
(Cold Patch for
a. Select a patch of ample size to extend at least
5/8-inch beyond the damaged area.
NOTE
NOTE
Surface coatings and surface refurbishing
kits will not repair leaks. Use repair kit
materials.
NOTE
When repairing de-ice boots and replacement
layers are being installed, exercise care to
prevent trapping air beneath the replacement
layers. If air blisters appear after material
is applied, they may be removed with a hypodermic needle. Should air blisters appear
after boots have been installed for a length
of time, it is permissible to cut a slit in the
de-ice boot, apply adhesive and repair in
accordance with the following cold patch repair procedures. An alternate method of
repair is to peel the de-ice boot back using
Toluol and reapply using 1300L cement.
a. Select a patch of ample size to cover damaged
area.
b. Clean area to be repaired with a cloth slightly
dampened with cleaner.
c. Buff area around damage with steel wool so that
area is moderately but completely roughened
d. Wipe buffed area clean with a cloth slightly dampened with cleaner to remove all loose particles,
e. Apply one even,thorough coat of 1300L cement
to the patch and to the corresponding damaged area
of the de-ice boot. Allow cement to set until it becomes tacky.
f. Apply patch to the de-ice boot with an edge or
the center adhering first, then work remainder of
patch down, being careful to avoid trapping air pockets.
g. Roll patch thoroughly with a stitcher roller, and
allow to set for ten or fifteen minutes.
h. Wipe patch and surrounding area from center of
patch outward with a cloth slightly dampened with
MEK.
i. Apply one light coat of A-56-B conductive cement
(B. F. Goodrich part number 74-451-11) to restore
conductivity.
If the correct size patch cannot be obtained,
one may be cut to the size desired from a
larger patch. If this is done, the edges
should be beveled by cutting with the shears
at an angle. These patches are manufactured
so they will stretch in one direction only. Be
sure to cut the patch selected so that the stretch
is in the width wise direction of the inflatable
tube.
b. Clean the area to be repaired with a cloth slightly
dampened with cleaner.
c. Buff the area around damage with steel wool so
that area is moderately but completely roughened.
d.. Wipe buffed area clean with a cloth slightly dampened with cleaner to remove all loose particles.
e. Apply one even, thorough coat of 1300L cement
to the patch and to the corresponding damaged area of
the de-ice boot. Allow cement to set until it becomes
tacky.
f. Apply patch to de-ice boot with the stretch in
the width-wise direction of the inflatable tubes, sticking edge of patch in place first, and working remainder down with a very slight pulling action so the rupture is closed. Use care not to trap air between
patch and de-ice boot.
g. Roll patch thoroughly with a stitcher roller and
allow to set for ten or fiteeen minutes.
h. Wipe patch and surrounding area, from the center of patch outward with a cloth slightly dampened
with cleaner.
I. Apply one light coat of A-56-B conductive cement
(B. F. Goodrich part number 74-451-11) to restore
conductivity.
NOTE
Satisfactory adhesion of patch to de-ice boot
should be reached in four hours; however, if
patch is allowed to cure for a minimum of
twenty minutes, de-ice boots may be inflated
to check the repair.
15-24B.
Area.)
APPROVED REPAIRS.
(Damage to Fillet
NOTE
NOTE
Satisfactory adhesion should be obtained in
four hours; however, if the patch is allowed
to cure for a minimum of twenty minutes,
the de-ice boots may be inflated to check the
repair.
15-24A.
Area.)
APPROVED REPAIRS.
(Damage to Tube
NOTE
This type of damage consists of cuts, tears
or ruptures to the inflatable tube area, and
a fabric-reinforced patch must be used.
This damage includes any tears or cuts to
the tapered area aft of the inflatable tubes.
a. Trim damaged area square and remove excess
material. Cut must be sharp and clean to permit a
good butt joint of the inlay.
b. Cut inlay from tapered fillet B. F. Goodrich
part number 74-451-21) to match cut out area.
c. Using Toluol, loosen edges of de-ice boot around
area approximately one and one-half inches from all
edges.
d. Clean area to be repaired with a cloth slightly
dampened with cleaner.
Revision 1
15-15
MODEL 210 & T210 SERIES SERVICE MANUAL
e. Lift back edges of cutout and apply one coat of
1300L cement to underneath side of loosened portion
of de-ice boot.
f. Apply one coat of 1300L cement to wing skin
underneath loosened edges of de-ice boot and extending one and one-half inches beyond edges of de-ice
boot into cutout area.
g. Apply second coat of 1300L cement to underneath
side of de-ice boot as outlined in step (e).
h. Apply one coat of 1300L cement to one side of a
two-inch wide neoprene-coated fabric tape (B. F.
Goodrich part number 74-451-22), allow to dry and
trim to size.
i. Reactivate cemented surfaces with Toluol and
apply reinforcing tape to wing skin, exercising care
to center tape under all edges of cutout.
j. Roll down tape on wing skin with stitcher
to assure good adhesion, being careful to avoid
creating air pockets.
k. Apply one coat of 1300L cement to top surface of
tape and allow to dry approximately five to ten min-
and one coat to veneer ply. Allow cement to set until
it becomes tacky.
g. Roll veneer ply to de-ice boot with a two-inch
rubber roller, applying a slight tension on veneer
ply when applying, to prevent trapping air.
h. Wipe patch and surrounding area from center
of patch outward with a cloth slightly dampened with
cleaner.
i. Apply one light coat of A-56-B conductive cement
(B. F. Goodrich part number 74-451-11) to restore
conductivity.
NOTE
B. F. Goodrich Repair Kit No. 74-451-C,
for repairing de-ice boots, is available
from Cessna Parts Distribution (CPD 2)
rollerthroughCessnaService Stations.
15-25. MATERIALS REQUIRED FOR INSTALLT
TION OF DE-ICE BOOTS
1. No. EC-1300L (EC-1403) Cement, Minnesota
Mining & Manufacturing Company.
utes.
l.
Reactivate cemented surfaces with toluol. Work2. Methyl-Isobutyl Ketone (MIBK).
3. Cleaning Solvent - Toluol.
ing toward cutout, roll down edges of loosened de-ice
4. Cleaning Solvent - Hexane.
boot, being careful to avoid creating air pockets.
Edges should overlap on tape approximately one inch.
5. Clean, lint-free cleaning cloths.
6.
Four yards clean, heavy canvas duck fabric
m. Roughen back surface of inlay repair material,
48 inches wide.
previously cut to size, clean with cleaner and apply
one coat of 1300L cement.
7. Several empty tin cans.
8. Three-inch paint brushes
n. Apply one coat of 1300L cement to wing skin in9. Two-inch rubber hand rollers.
side of cutout area and allow to dry.
10. 1/4-inch metal hand stitcher roller, B. F.
o. Apply second coat of 1300L cement to back side
Goodrich Company (Part Number 3306-10).
of inlay material and allow to dry.
11. Carpenters' chalk line.
p. Reactivate cemented surfaces with Toluol and
12. One-inch marking tape.
carefully insert inlay material with feathered edge
aft. Working from wing leading edge aft, roll down
13. Steel measuring tape.
14. Sharp knives.
inlay material carefully to avoid trapping air.
15. Fine sharpening stone.
q. Roughen area on outer surface of de-ice boot
16. No. EC-539 Sealing Compound, Minnesota
and inlay with steel wool, one and one-half inches
Mining & Manufacturing Company.
on each side of splice. Clean with cleaner and ap17. No. A-56-B Cement, B. F. Goodrich Comply one coat of 1300L cement to this area.
pany (Part Number 3306-15).
r. Apply one coat of 1300L cement to one side of
two-inch wide neoprene-coated fabric tape, trim to
18. GACO-700-A Coating, Gates Engineering Co.,
Delaware 19899.
size and center tape over splice on all three sides.Wilmington,
s. Roll down tape on de-ice boot with stitcher
15-26. REPLACEMENT OF DE-ICE BOOTS. To
roller to assure good adhesion, being careful to
remove or loosen installed de-ice boots, use toluol
avoid creating air pockets.
t. Apply one light coat of A-56-B conductive
or toluene to soften the "cement" line. Apply a
minimum amount of this solvent to the cement line
cement (B. F. Goodrich part number 74-451-11) to
as tension is applied to peel back the boot. Removal
restore conductivity.
should be slow enough to allow the solvent to undercut the cement so that parts will not be damaged. To
15-24C. APPROVED REPAIRS. (Damaged Veneer,
install a wing de-icer boot, proceed as follows:
loose from De-ice Boot.)
Clean the metal surfaces and the bottom side of
a. Peel and trim loose veneer to the point where
adhesion of veneer to de-ice boot is good.
the de-icer thoroughly with Methyl Ethyl Ketone or
Methyl Isobutal Ketone. This shall be done by wiping
b. Roughen area in which veneer is removed, with
the surfaces with a clean, lint-free rag soaked with
steel wool, rubbing parallel to cut edge of veneer
ply to prevent loosening it.
the solvent and then wiping dry with a clean, dry,
lint-free rag before the solvent has time to dry.
c. Taper edges of veneer down to tan rubber ply by
Place one inch masking tape on wing to mask off
rubbing parallel to edges with steel wool and MEK.b.
boot area allowing ½ inch margin. Take care to mask
d. Cut a piece of veneer material (B. F. Goodrich
so that clean-up time will be reduced.
andaccurately area
part number 74-451-23) to cover damaged
StirEC-1300L cement thoroughly before using.
extend at least one-inch beyond, in all directions.
Brush one even, light coat onto leading edge and to
e. Mask off an area one-half inch larger in length
and width than size of veneer patch
rough side of boot, brushing well into rubber. Allow
Apply-one
of 1300L
coat cement to damag ed area,
cement to air dry until cement does not transfer to
15-16
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
fingers when touched.
Then apply a second coat to
each of the surfaces and allow to dry. Apply a
vacuum to the boots when they are installed to help
smooth out wrinkles.
d. Place a straight line along the leading edge line
and a corresponding line on the inside of the de-icer
boot if it does not have a centerline. Securely attach
hoses to de-icer connections. Position centerline of
boot with leading edge line, using a clean, lint-free
cloth, heavily moistened with toluol, reactivate surface of cement on wing and the boot in small, spanwise areas approximately 6-inches wide. Avoid excessive rubbing of cement, which would remove it
from the surface of the wing. Utilize enough help to
hold boot steady during installation, and caution them
against handling cemented surfaces. Roll boot firmly
against leading edge, being careful not to trap any air
between boot and leading edge surface. Always roll
parallel to the inflatable tubes. Should the boot attach
"off course", pull it up immediately with a quick
motion, and reposition properly. Avoid twisting or
sharp bending of boot. Finally, roll the entire
surface of the boot parallel to tubes, applying pressure. Use the metal stitcher roller between tubes
and around connections. Should an air pocket be
encountered, carefully insert a hypodermic needle
and allow air to escape. Do not puncture the inflatable tubes at any time. Fill any gaps between
adjoining boots with GACO N-700-A Neoprene coating
(Gates Engineering Co., Wllmington, Delaware
19899). Apply a coat of the Neoprene coating along
trailing edge of boot to the surface of the skin to form
a neat, straight fillet.
a. Remove masking tape and clean surfaces with
tohsol
15-26A. WING AND HORIZONTAL STABILIZER
THREE-CYCLE DE-ICE SYSTEM. (Beginning
with 21062969.) (See figure 15-5A.)
15-26B. DSCRIPTION. The system consists of
pneumatically-operated boots, an engine-driven
pneumatic pump, an annunciator light to monitor
system operation, system controls and the hardware
necessary to complete the system.
15-26C. SYSTEM OPERATION. The boots expand
and contract, using pressure or vacuum from the
engine-driven vacuum pump. Normally, vacuum is
applied to all boots to hold them against the leading
edge surfaces. When a de-icing cycle is initiated,
the vacuum is removed and a pressure is applied to
"blow up," the boots. Ice on the boots will then be
removed by normal in-flight air forces. Controls
for the system consist of a spring-loaded on-off
rocker switch on the left switch and control panel, a
pressure indicator light on the upper left side of the
instrument panel, and a 5-amp "pull-off" type circuit
breaker on the left sidewall circuit breaker paneL
The two-position de-icing switch, labeled DE-ICE
PRESS, is spring-loaded to the normal off (lower)
position. When pushed to the ON (upper) position and
released, it will activate one de-icing cycle. Each
time a cycle is desired, the switch must be pushed to
the ON position and released. If necessary, the
system can be stopped at any point in the cycle (de-
flating the boots) by pulling out the circuit breaker
labeled WING, DE-ICE. During a normal de-icing
cycle, the boots will inflate according to the following
sequence: first, the horizontal stabilizer boots will
inflate for approximately six seconds, then the inboard boots inflate for the next six seconds, followed
by the outboard wing boots for another six seconds.
The total time required for one cycle is approximately
18 seconds. The pressure indicator light, labeled
DE-ICE PRESSURE, should illuminate when the horizontal stabilizer boots reach proper operating pressure. At lower altitudes, it should come on within
one to two seconds after the cycle is initiated and remain on for approximately 17 seconds if the system
is operating properly. At higher altitudes, the light
will come on initially within three seconds and will go
off for one to three seconds during sequencing. The
system may be recycled six seconds after the light
goes out. The absence of illumination during any one
of the three sequences of a cycle indicates insufficient
pressure for proper boot inflation and effective deicing ability. An ice detector light is also installed
to facilitate detection of wing ice at night or during
reduced visibility. The ice detector light system
consists of a light installed on the left side of the
cowl deck forward of the windshield which is positioned to illuminate the leading edge of the wing,
and a rocker-type switch, labeled DE-ICE LIGHT,
located on the left switch and control panel.
15-26D. FLIGHT INTO KNOWN ICING EQUIPMENT
AND SYSTEMS. (Beginning with serial 21063253.)
(See figure 15-5B.)
15-26E. DESCRIPTION. A night into known icing
equipment package may be installed on the airplane.
For operations in known icing conditions as defined
by the FAA, the following Cessna (drawing number
1200254) and FAA approved equipment must be installed and operational:
1. Wing horizontal stabilizer and vertical fin
leading edge pneumatic de-ice boots.
2. Propeller anti-ice boots.
3. Windshield anti-ice panel.
4. Heated pitot tube (high capacity).
5. Heated stall warning transducer (high
capacity).
6. Ice detector light.
7. 95-amp alternators. (Thru 1982
models).
8. Dual 60-amp alternators. (Beginning with
1983 models).
9. Control surface static dischargers.
10. High capacity vacuum pump (thru 1981 models).
11. Dual vacuum pumps. (Beginning with 1982
models).
Service information on this equipment when installed
on known icing certified aircraft is contained in the
following paragraphs.
15-26F. WING, HORIZONTAL STABILIZER AND
VERTICAL FIN DE-ICE SYSTEM. (Beginning with
serials 21063253.) (See figures 15-5C and 15-5D. )
Revision 1
15-17
MODEL 210 & T210 SERIES SERVICE MANUAL
Detail C
C
A
1. Wing De-Ice Boot
2. De-Ice Pressure Light
3. Switch and Circuit
Breaker Panel
4. Vacuum Pump
5. Pressure Valve
8
1
10
7. Cross
Figure
15-5A. Wing and Horizontal Stabilizer Three-Cycle De-Ice SystemDetail
15-18
Revision
1
Switch
l
8. Pressure
9. Stabilizer De-Ice Boot
10. Flow Control Valve
11. Grommet
12. Tube Assembly
17
16
11
13. Cover Assembly 18
15
14.
Lens
18.
De-Ice Bulb
BEGINNING WITH 21062969
Detail A
B
11
MODEL 210 & T210 SERIES SERVICE MANUAL
*
._
MODEL 210 & T210 SERIES SERVICE MANUAL
21063641 THRU 21064535
6 21 22 23 24 25
21064536 THRU 21064772
MODEL 210 & T210 SERIES SERVICE MANUAL
BEGINNING WITH 1983 MODELS
SERIAL 21064773 & ON
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36. Lo-Vacuum Warning Light
37. Heated Stall Warning Circuit Breaker
38. Stall Warning Heat Switch
Figure 15-5B.
Known Icing Equipment Installation (Sheet 3 of 3)
Revision 1
15-21
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
15-26G. DESCRIPTION. The system consists of an
engine-driven vacuum pump, pressure control valve,
relief valve adjustment should be maintained in accordance with procedures outlined in the applicable
paragraph in Section 16 of this manual. If the vacuum
relief valve is set too low, suction to the gyros will
drop momentarily during the boot inflation cycle.
This suction variation can be corrected with proper
vacuum relief valve adjustment. The standard vacuum pump is replaced with a larger capacity vacuum
vacuum relief valve, flow control valves, pressure
pump. Beginning with 1982 models dual vacuum
switch, timer and boots mounted on the leading edge
of each wing, horizontal stabilizer and the vertical
fin. The aircraft vacuum system components also
serve the de-ice vacuum system, and the vacuum
pumps and dual control valves are components of the
system. An ice detector light is incorporated in the
left side of the cowl deck below the windshield to aid
in checking for ice formations during night operation.
A few aircraft which are not certified for
flight into known icing conditions may have
this system installed.
15-26H.
TROUBLE SHOOTING -- WING, HORIZONTAL STABILIZER AND VERTICAL-IN DE-ICE SYSTEM.
AOMD
*~~~~~~~
#SE83-12.
isequipped
The
system
DESCRIPTION.
(Figure1-C
with
de-ice flow valves15-26K.
three
to the
vacuum pimps
2 de-ice boots.
Revision
15-22
15-26J. DE-ICE FLOW VALVE. (Serial 21062969
thru 21064802.) B. F. Goodrich part number 3D235701.
15-26L. DE-ICE FLOW VALVE OVERHAUL. If it
becomes necessary to overhaul a de-ice flow valve
follow-the
(B. F. Goodrich part
~ number 3S2357-01),
~
procedures outlined in Service Information Letter
SE83-12.
15-22 Revision--2
15-26K. DESCRIPTION. The system is equipped
with three de-ice flow valves (Figure 15-5C, items
13 and 25.) The valves are electrical solenoid
operated and route pressure and vacuum from the
'-
MODEL 210 & T210 SERIES SERVICE MANUAL
Beginning with 21064536
L
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Fuselage Skin
Lens
A
Cover Assembly
Hose
De-Ice Control Valve
Line
11. Grommet
This installation is a required
r' 1
component for flight in known
lcing certified aircraft.
.fy~~~~~~~
7DelC
,8~~~~~~~~~~~~~~~~~~8
1. Bulb
2.
3.
4.
5.
6.
7.
8.
9.
10.
sNOTE
7
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11
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Figure 15-5D. Wing, Horizontal Stabilizer and Vertical Fin De-Ice System (Sheet 1 of 4)
Revision 1
15-25
MODEL 210 & T210 SERIES SERVICE MANUAL
26
MODEL 210 & T210 SERIES SERVICE MANUAL
9.
15-28
Lin
'"
Figure 115-5D.
Revision
Wing, Horizontal Stabilizer and Vertical Fin De-Ice System (Sheet 4 of 4)
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10. De-Ice Pressure
Control Valve
28. Vacuum Pump
29. Check Valves
o
^O°°°°
nin
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'
NaI E
Dual Vacuum Pumps (28) beginning with 1982 Models.
Dual Pressure Control Valves (9) beginning with 1983
°
MODEL 210 & T210 SERIES SERVICE MANUAL
15-26M. DE-ICE SYSTEM FUNCTIONAL CHECK
(KNOWN ICING). (See figure 15-5E.)
a. Electrical Controls Check:
1. Check wing de-ice circuit breaker is closed.
2. Check de-ice pressure switch is off (springloaded to off position).
3. Turn master switch on.
4. Press de-ice pressure light to check light
circuit and bulb. Make sure dimming shutter is open.
5. Turn master switch off.
b. Vacuum Relief Valve(s) Adjustment.
1. Refer to Section 16 of this manual for vacuum
relief valve(s) adjustment.
c. Preflight System Check:
1. With vacuum relief valve(s) adjusted and engine running from 2200 to 2500 rpm, check both
buttons on the suction gage are retracted out of sight
and vacuum is normal.
2. Place de-ice pressure switch on and release.
3. Check that de-ice pressure light comes on
within one second, remains on for 18 seconds, then
off.
4. Check boots for inflation during 18 second
cycle as follows: first six seconds tail section boots,
then inboard wing boots for next six seconds, finally
the outboard wing boots inflate for six seconds completing one cycle.
5. The absence of or slow illumination of the
de-ice pressure light during any one of the three
sequences of a cycle indicates insufficient pressure
for proper system operation,
d. Timer Check:
1. Refer to paragraph 15-26U for timer check.
e. Air Pressure Check (See figure 15-5E):
NOTE
This check may be performed in the engine
1. Disconnect both pump pressure hoses (8) from
vacuum pumps (1).
Connect a source of clean regulated dry air
pressure (21 ±1 psig) fitted with a hand-operated
valve or check valve and an in-line air pressure
gauge to right pump pressure hose (8).
NOTE
A test kit (No. 343) for testing vacuum and
pneumatic de-ice system is available from
Airborne, 711 Taylor Street, Elyria, Ohio
44035, or from Cessna Parts Distribution (CPD
2) through Cessna Service Stations. This kit
contains the necessary equipment and
supplemental instructions to perform this check.
3. Disconnect left and right vacuum inlet hoses
from left and right vacuum pumps (1).
4. Disconnect electrical leads from pressure
control valves (3).
CAUTION
Do not attempt air pressure check with
de-ice timer module connected into the
circuit.
5. Connect a vacuumsource (5.6 in. Hgminimum)
to right pump vacuum hose.
6. Connect a switched 28VDC electrical source to
right pressure control valve (3).
7. Insert pressure probe equipped with vacuum
pressure gage into the rubber hose connecting tail
boots with tail boot flow valve.
8. Turn on pressure and vacuum sources.
Verify that pressure flow is being vented overboard
at right pressure control valve and no flow is present
either in or out of disconnected hoses at left vacuum
pump. Pressure gage on probe should read 4. 5-4. 6
in. Hg vacuum.
9. Switch on electrical power to right pressure
control valve and actuate tail boot flow control manually.
NOTE
Flow valves can be actuated mechanically
by depressing the solenoid plunger inward
using the fingers. This procedure eliminates the necessity of disconnecting and
reconnecting electrical leads.
10. Overboard flow at pressure control valve
should stop and pressure air should inflate tail
boots. Pressure gage should show 18 ±. 5 psi with
audible venting of pressure air from pressure regulator valve (7) evident. Recheck for absence of
airflow out of left pressure control valve.
11. With pressure control valve energized turn
off pressure source using hand-operated valve.
Pressure leak-down as shown by probe pressure
gage should be 2 psi per minute or less. Use soap
and water solution to locate leaks, turn off power
to left pressure control valve, repair leaks and
restest until leak-down rate is within tolerance.
12. Insert pressure probe into hose connecting
outboard wing boots with outboard boot flow control
valve and repeat steps 8 thru 11 noting leaks.
13. Insert pressure probe into hose connecting
inboard wing boots with inboard boot flow control
valve and repeat steps 8 thru 11 noting leaks.
14. Disconnect pressure and vacuum sources
from right vacuum pump hoses and connect to left
pump hoses.
15. Turn on pressure and vacuum sources.
Verify that pressure flow is being vented overboard
at left pressure control valve and no flow is present
either in or out of disconnected hoses at right pump.
Probe pressure gauge should read 4. 5-5. 6 in. Hg
vacuum.
16. Switch on electrical power to left pressure
control valve. Overboard flow at pressure control
valve should stop. Check for no airflow from right
pressure control valve and audible venting of pressure
air from pressure regulator valve (7) evident.
17. With probe air pressure gauge inserted into
hose connecting any flow valve with its associated
de-ice boot, actuate flow valve manually, and recheck
probe air pressure gauge reads 18 ±. 5 psi.
18. Disconnect test equipment and reconnect pressure and vacuum lines to vacuum pumps.
19. Reconnect wiring to pressure control valves.
Revision 3
15-31
MODEL 210 & T210 SERIES SERVICE MANUAL
15-26N. DE-ICE BOOT REPAIR. (COLD PATCH.)
Follow procedures outlined in paragraph 15-24.
f. Timer output shall complete the cycle then shut
off all outputs.
15-26P. DE-ICE BOOT TYPES OF DAMAGE AND
REPAIR. Follow procedures outlined in paragraphs
15-24A, 15-24B, and 15-24C.
NOTE
15-26Q.
MATERIALS REQUIRED FOR INSTALLA-
TION OF DE-ICE BOOTS.
in paragraph 15-25.
Use the materials listed
15-26R. REPLACEMENT OF DE-ICE BOOTS. Follow the procedures outlined in paragraph 15-26.
15-26S.
TIMER (See figure 15-5E.)
15-26T. DESCRIPTION. The timer, located on the
underside of the glove box, controls the time the deice boots are inflated.
Do not check voltage levels without a load
attached; readings may be erroneous.
15-26V. FUNCTIONAL TEST OF TIMER. (1983
Models and on) (See figure 15-5E, Sheet 2.)
a. Connect timer as shown in wiring schematic.
b. Set the voltage at 28 VDC and turn the control
switch on.
c. Record the time each light is on.
d. The recorded times shall be as shown in the
chart (sheet 2) *10% at 28 VDC.
e. Turn control switch on, then release to off.
f. The timer output shall complete the cycle and
then shut off all outputs.
15-26U. FUNCTIONAL TEST OF TIMER. (Thru 1982
Models) (See figure 15-5E, Sheet 1.)
a. Connect timer as shown in the wiring schematic.
b. Set voltage at 28 VDC, and turn control switch on.
Record the time each light is on.
The recorded times shall be as shown in the
c.
d.
chart * 10% at 28 VDC.
e. Turn control switch on, then release to off.
NOTE
Do not check voltage levels without a load
attached; readings may be erroneous.
g. Vary the voltage from 22-31 VDC and repeat
step f. Timer must continue to operate at these
voltages within the time frame shown in chart.
NOTE
BLACK
TIMING CHART
*
*
BLUE
YELLOW
WRITE/BLUE
.1 AMP LAMP
20 RESISTOR, 50 WATT OR 24 - 32 VDC SOLENOD
Figure 15-5E.
15-32
VIOLET
18 SECONDS
6 SEC.
6SEC.
SEC.
Wing. Horizontal Stabilizer and Vertical Fin De-Ice System Timer Test (Sheet 1 of 2)
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
BLACK
ORANGE
MS 35058-30
CONTROL SWITCH
r
GREEN
TIMER
28 VDC
BLUE
YELLOW
NOTE
WHITE/BLUE
Black wire is the ground wire.
VIOLET
The unit shall, once control
switch is activated, complete
one cycle.
*
*
*
Reactivation of the
control switch during the initial cycle shall not interrupt
or reset.unit until one cycle
is completed.
* .1 AMP LAMP, 28 VDC
20 OHM RESISTOR, 50 WATT OR 24-32 VDC SOLENOID (20 OHM)
O 65 OHM RESISTOR, 10 WATT OR 24-32 VDC SOLENOID (65 OHM)
TIMING CHART
WIRE COLOR
BLUE
YELLOW
WHITE/BLUE
VIOLET
GREEN
Figure 15-5E.
TIME ON (SECONDS)
18 SECONDS
6 SEC.
_
6 SEC.
6 SEC.
18 SECONDS
Wing, Horizontal Stabilizer and Vertical Fin De-Ice System Timer Test (Sheet 2 of 2)
15-26W. PROPELLER ANTI-ICE BOOTS (KNOWN
ICING EQUIPMENT). Aircraft certified for night
into known icing conditions must have propeller antiice boots installed and operational. Refer to paragraph 15-27 for this installation.
15-26X. WINDSHIELD ANTI-ICE PANEL (KNOWN
ICING EQUIPMENT). Aircraft certified for flight
into known icing conditions must have a windshield
anti-ice panel installed and operational. Refer to
paragraph 15-32D for this installation.
Beginning with 1983 models, separate switches,
labeled PITOT HEAT and STALL HEAT, on the left
switch and control panel operate the heaters. Two
10-amp "push-to-reset" type circuit breakers,
labeled PITOT HEAT and STALL HEAT, on the left
sidewall circuit breaker panel protect the systems.
When the aircraft is on the ground, a resistor is
introduced into the stall warning heater circuit by
the nose wheel squat switch in order to prevent oveheating.
15-26AA.
15-26Y. PITOT TUBE AND STALL WARNING
HEATERS. (KNOWN ICING) (See figure 15-5B.)
15-26AB.
15-26A. DESCRIPTION. A special pitot tube with a
larger inlet and a higher capacity heating element
and a higher capacity heated stall warning transducer
are installed in the left wing on aircraft certified for
flight into known icing conditions. These systems
assure proper airspeed indications and stall warning
in the event icing conditions are encountered. They
are designed to prevent ice formation rather than
remove it once formed. Thru 1982 models both
systems are controlled by a rocker switch, labeled
PITOT HEAT, on the left switch and control panel.
REMOVAL AND INSTALLATION.
(See
Section 17.)
ICE DETECTOR LIGHT.
15-26C. DESCRIPTION. An ice detector light is
flush-mounted on the left side of the cowl deck to
facilitate the detection of wing ice at night or during
reduced visibility by lighting the leading edge of the
wing. Components of the system include the ice
detector light, a two-position rocker-type switch.
labeled DE-ICE LIGHT, on the left switch and a 5amp "push-to-reset" type circuit breaker, labeled
CABIN LIGHTS on the left sidewall circuit breaker
panel. The richer switch is spring-loaded to the
Revision 2
15-33
MODEL 210 & T210 SERIES SERVICE MANUAL
off (lower) position and must be held in the ON (upper)
position to keep the ice detector light illuminated.
15-26AD. 95-AMP ALTERNATOR INSTALLATION.
(thru 1982 Models) (See Section 17.)
15-26AE. DUAL 60-AMP ALTERNATOR INSTALLATION. (Beginning with 1983 Models.) To provide
electrical system redundancy dual 60-amp alternators
must be installed and fully operational on aircraft
certified for flight into known icing conditions. See
Section 17.
15-26AF.
CONTROL SURFACE DISCHARGERS.
15-26AG. DESCRIPTION. Wick type static dischargers may be installed on the trailing edge surfaces of
the ailerons, elevators and rudder of the aircraft.
One type discharger is fabricated with the wick and
base combined into an integral unit; in the other type,
the wick is attached to the base by a threaded fitting,
and may be replaced without removing the base from
the aircraft. The installation of static dischargers
reduces the build-up of static electricity on the airframe as a consequence of flying through haze, dust,
rain, snow or ice crystals. In some cases, if dischargers are not installed or not functioning as a
result of age or repeated exposure static electricity,
static build-up can result in the loss of usable radio
signals on all communication and navigation equipment. Whenever static dischargers are installed,
replaced, and at regular intervals during their service life, resistance checks should be performed to
determine their effectiveness in reducing static
build-up.
15-34
Revision 2
15-26AH. RESISTANCE CHECK. Since static dischargers lose their effectiveness with age and exposure to static electricity, they should be checked
with a 500 to 1000 volt capacity megohmmeter every
500 hours or annually; whichever occurs first. Megohmmeters may be purchased from the following
source:
James G. Biddle Co.
Plymouth Meeting, PA 19462
NOTE
A GOOD aircraft ground must be established
in order to perform RELIABLE resistance
checks on the control surface dischargers.
Perform the following resistance checks on each control surface discharger and replace those which do
not conform to the resistance requirements.
a. If the wick and base of the discharger are an
integral unit, the resistance from the base of the
discharger to a good aircraft ground should check
2. 5 milliohms maximum.
b. If the wick can be separated from the base, the
resistance from the base to a good aircraft ground
should check 1.0 ohm maximum.
c. Connect the EARTH terminal to the base of the
discharger and check the resistance at the tip of the
wick. The resistance should check 1 to 100 megohms
for both types of dischargers.
WARNING
So not bend the wick during the preceding
check, since wicks have a higher resistance
when bent.
MODEL 210 & T210 SERIES SERVICE MANUAL
15-27. PROPELLER ANTI-ICE SYSTEM. The system is of an electrothermal type, consisting of
electrically-heated de-ice boots bonded to each
propeller blade, a slip ring assembly for power
distribution to the propeller de-ice boots, a brush
block assembly to transfer electrical power to the
rotating slip ring, and a timer to cycle electrical
power to the de-ice boots in proper sequence. A
rocker switch labeled PROP A/ICE, located on the
pilot's lower left-hand panel, controls the propeller
de-ice system. A circuit breaker labeled PROP
A/ICE, located in the left circuit breaker panel,
protects the propeller de-ice system. A propeller
de-ice ammeter, located on the upper left instrument
panel, indicates amperage for the propeller de-ice
system.
The de-ice system applies heat to the surfaces of the
propeller blades where ice would normally adhere.
This heat, plus centrifugal force and the blast from
the airstream, removes accumulated ice. Each deice boot has two separate electrothermal heating ele15-27A.
TROUBLE SHOOTING ---
TROUBLE
ELEMENTS DO NOT HEAT.
ments, and inboard and an outboard section. Each
boot has three leads extending from a tab at the bottom of the boot. Each electrical lead is identified by
a letter. The letter "G" stands for ground. The
letter 'T' stands for inboard, and the letter "O" stand
for outboard. When the PROP A/ICE switch is turned
on, the timer provides power through the brush block
and slip ring to the outboard element of the propeller
for approximately 20 seconds ±1 second. The timer
then switches power to the inboard element of the
propeller for approximately 20 seconds *1 second.
The complete cycle is then repeated. This outboardinboard sequence is very important since the loosened
ice, through centrifugal force, moves outboard. Heat
ing may begin at any phase in the cycle, depending on
timer position when the switch was turned off from
previous use. Ground checkout of the system is permitted with the engine not running. Propeller remova
is necessary before propeller de-ice system components, except for the brush block assembly, timer,
ammeter, circuit breaker and switch can be removed
or installed.
PROPELLER ANTI-ICE SYSTEM.
PROBABLE CAUSE
Circuit breaker out or defective.
REMEDY
Reset circuit breaker. If it pops out
again, determine cause and correct.
Replace defective parts.
Defective wiring.
Repair or replace wiring.
Defective switch.
Replace switch.
Defective timer.
Replace timer.
Revision 2
15-34A/(15-34B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
15-27A. TROUBLE SHOOTING --- PROPELLER DE-ICE SYSTEM (Cont).
PROBABLE CAUSE
TROUBLE
REMEDY
ELEMENTS DO NOT HEAT.
Defective brush-to-slip ring
connection.
Check alignment. Replace defective
parts.
SOME ELEMENTS DO NOT
HEAT.
Incorrect wiring.
Correct wiring.
Defective wiring.
Repair or replace wiring.
Defective timer.
Replace timer.
Defective brush-to-slip ring
connection.
Check alignment.
parts.
Defective element.
Replace element.
CYCLING SEQUENCE NOT
CORRECT OR NO CYCLING.
Crossed connections.
Correct wiring.
Defective timer.
Replace timer.
RAPID BRUSH WEAR,
FREQUENT BREAKAGE,
SCREECHING OR
CHATTERING.
Brush block or slip ring out of
alignment.
Align properly.
15-27B. SLIP RING REMOVAL.
(See figure 15-6.)
WARNIN
Be certain magneto is grounded before
turning propeller.
a. Remove spinner attaching screws (22) and remove spinner (12), spinner support (20) and spacers
(21). Retain spacers (21).
b. Remove engine cowling as required for access
to propeller mounting nuts (24) and washers (23).
c. Loosen all propeller mounting nuts (24) approximately 1/4-inch and pull propeller forward until
stopped by mounting nuts (24).
NOTE
from engine crankpropeller
is
Asseparated
As propeller is separated from engine crankshaft fange, oil will drain from propeller and
engine cavities.
CAUTION
Use caution when removing propeller. Removing propeller without the de-ice slip
the spinner bulkhead, since the slip ring is
mounted to the bulkhead. Wires should be
identified according to wiring diagrams to
facilitate reassembly. During removal,
installation or other maintenance, use care
to prevent damaging slip ring and brushes.
d. Remove safety wire and loosen clamps (13).
e.
Replace defective
Remove nuts, washers, de-ice lead wires and
head (7). Tag lead wires to facilitate reinstallation.
f. Remove all propeller mounting nuts (24) and
washers (23) and pull propeller forward to remove
from engine crankshaft (25).
g. Remove slip ring (6).
15-27C. SLIP RING INSTALLATION. (See figure
15-6.)
a. Install slip ring (6) and aft spinner bulkhead (7).
b. Install de-ice boot lead wires and slip ring lead
wires, screws, washers and nuts in aft spinner bulkhead (7).
c. Install propeller and install washers (23) and
propeller mounting nuts (24).
d. Secure aft spinner bulkhead (7) to propeller with
screws.
e. Tighten propeller mounting nuts to a torque of
55 to 60 lb. ft.
f. Tighten clamps (13) with clamp screw housings
180 ° apart to maintain balance. Safety wire clamp
screw housings to clamps as shown in view B-B.
g. Install spacer (21) and spinner support (20) in
spinner (12) and install spinner on propeller.
15-2. SLIP RING ALIGNMENT CHECK. After installation, slip ring must be checked for run-out.
NOTE
Excessive slip ring run-out will result in
severe arcing between slip ring and brushes,
and cause rapid brush wear. If allowed to
continue, this condition will result in rapid
deterioration of slip ring and brush contact
Revision 1
15-35
MODEL 210 & T210 SERIES SERVICE MANUAL
10
Restrainer strap to start
at point "A." Wrap restrainer strap clockwise.
End strap at point "B."
Trim strap length as
necessary.
10.
Boot
11.
Restraining Strap .10-ich
1.00 ± .10-inch
1. 250-inch
THREADLESS PROPELLERS
BEGINNING WITH 21061574
10
.38-inch
Restrainer strap to start
in this area (approximately
120° from lead strap. Wrap
around prop blade twice so
a double thickness will cover
120°
1.38-inch
the de-ice lead strap. Trim
restrainer strap so it will
end approximately as shown.
.19-inch
0.12-inch (approx.)
Figure 15-6.
Propeller Anti-Ice System (Sheet 3 of 3)
surfaces, and lead to the eventual failure of
the propeller de-icing system.
a. Securely attach a dial indicator gage to the engine and place the pointer on the slip ring.
b. Rotate the propeller slowly by hand, noting the
deviation of the slip ring from a true plane as indicated on the gage.
c. Check that the total run-out does not exceed 0.010
inch (± 0.005 inch), and that the total is not exceeded
within any four inches of slip ring travel.
NOTE
Care must be taken to exert a uniform push
or pull on the propeller to avoid a considerable error in the readings caused by loose
15-38
Revision 1
fitting thrust bearings.
If slip ring run-out is within the limits specified,
no corrective action is required. If the run-out is not
within limits specified, the slip ring will have to be
removed and returned to the claims department of
Cessna Supply Division, and a new part ordered.
15-29. REMOVAL OF PROPELLER ANTI-ICE
TIMER. (See figure 15-6.)
a. Ensure that aircraft electrical power is off and
PROP A/ICE circuit breaker is pulled.
b. Gain access to left elevator control torque tube
support (19), forward of left instrument panel.
c. Remove screws, timer (2) and spacers (17) from
nutplates.
MODEL 210 & T210 SERIES SERVICE MANUAL
15-29A. INSTALLATION OF PROPELLER ANTIICE TIMER. (See figure 15-6.)
a. Install spacers (17), timer (2) and screws in
nutplates in left elevator control torque tube support
(19).
b. Push in PROPA/ICE circuit breaker.
15-29B. PROPELLER ANTI-ICE SYSTEM
AMMETER. (See figure 15-6.)
15-29C. DESCRIPTION. An ammeter is utilized in
the propeller anti-ice system to visually monitor the
amperage being applied to that system.
15-29D. REMOVAL. (See figure 15-6.)
a. Ensure that aircraft electrical power is off and
15-29F.
PROP A/ICE circuit breaker is pulled.
b. Gain access to forward side of instrument panel
(right side thru 21062273; left side beginning with
21062274).
c. Unscrew bezel (26) and remove along with O-ring
(27).
d. Remove body (29) forward out of instrument
panel (28).
15-29E. INSTALLATION. (See figure 15-6.)
a. Install body (29) aft through hole in instrument
panel- (28):
b. Install O-ring (27) and screw bezel (26) on
threads of body (29).
c. Push in PROP A/ICE circuit breaker.
TROUBLE SHOOTING -- PROPELLER ANTI-ICE SYSTEM AMMETER.
REMEDY
PROBABLE CAUSE
TROUBLE
AMMETER READING
BELOW GREEN ARC.
Open anti-ice boot element.
Replace boot.
AMMETER READING
ABOVE GREEN ARC.
Shorted anti-ice boot element.
Replace boot.
NO AMMETER READING.
(Boots are heating)
Faulty ammeter shunt
Replace ammeter shunt.
Open circuits in wiring to
ammeter.
Repair wiring.
Faulty ammeter.
Replace ammeter.
Faulty system component.
Determine cause and correct.
NO AMMETER READING.
(Boots not beating)
15-30. TIMER TEST.
a. Remove connector plug of wire harness from timer
and jump power input socket of wire harness to timer
Timer P/N
3E1540-1
C165020-0101
PowerInput Pin& Socket
B (14VDC)
B (28VDC) (24-32)
Ground Pin
A (14VDC)
G (28VDC)
b. Jump timer ground pin to ground.
c. Turn on De-Icing System.
d. Check timer operation per the chart preceding
step "b." (Use a voltmeter.)
e. Check volts to ground in each case. If engine is
not running, and auxiliary power is not used, voltage
will be battery voltage and cycle time may be slightly
longer than indicated.
f. Hold voltmeter probe on the pin until the voltage
input pins. (Refer to chart following this step for pin
identification.)
OutputSequence, Time, Voltage
C, D 34 seconds each
C, D 20 seconds each
Time Repeat Cycle
Time (sec)
74
40
drops to 0. Move the probe to the next pin in the sequence shown in the chart. Check voltage at each pin
in sequence. When correctness of the cycling sequence
is established, turn propeller De-Icing switch off at
the beginning of one of the on-time periods, and record the letter of the pin at which the voltage supply
is present.
Revision 1
15-39
MODEL 210 & T210 SERIES SERVICE MANUAL
c. Draw a line on the centerline of the leading edge
of the blade. Position the pattern centerline over the
leading edge centerline. Position pattern so bottom
of boot is 1/2" below spinner cutout. Draw a line on
the propeller hub on each side of the pattern boot
strap where it crosses the hub. Check boot strap
position by fitting restraining strap on the hub and
comparing its position with the marked position of the
strap.
d. Mask off an area 1/2" from each side and outer
end of the pattern, and remove the pattern.
e. Mix EC-1300L cement (Minnesota Mining & Mfg.
Co.) thoroughly. Surfaces shall be above 60- F (15*
Centigrade) prior to applying cement. During periods
of high humidity, care shall be taken to prevent moisture condensation due to the cooling effect of the evaporating solvent. This can be done by warming the
area with a heat gun or heat lamp. Apply one even
brush coat of EC-1300L cement to the cleaned metal
surface. Allow to air dry for a minimum of one hour,
then apply a second even brush coat of EC-1300L cement.
f. Moisten a clean cloth with Methyl Ethyl Ketone
and clean the unglazed back surface of the boot,
changing cloths frequently to avoid contamination of
the cleaned area.
J Apply one even coat of EC-1300L cement to back
surface of boot. It is not necessary to cement more
than 1/2 of the boot strap.
h. Using a silver-colored pencil, mark a centerline
along the leading edge of the propeller blade and a
corresponding centerline on the cemented side of the
boot.
i. Reactivate the surface of the cement using a
clean, link-free cloth, heavily moistened with toluol.
Avoid excessive rubbing of cement, which would remove the cement.
j. Position the boot centerline on the propeller
leading edge, starting at the hub end at the position
marked. Make sure that boot strap will fall in the
position marked. Tack the boot centerline to the
leading edge of the propeller blade. If the boot is
allowed to get off-center, pull up with a quick motion
and replace properly. Roll firmly along centerline
with a rubber roller.
k. Gradually titing the roller, -work the boot carefully over either side of the blade contour to avoid
trapping air in pockets.
1. Roll outwardly from the centerline to the edges.
If excess material at the edges tends to form wrin
kles, work them out smoothly and carefully with
fingers.
m. Apply one even coat of EC-539 (Minnesota Mining & Mfg. Co.), mixed per manufacturer's instructions, around the edges of the installed boot.
n. Remove masking tape from the propeller and
clean the surface of the propeller by wiping with a
clean cloth dampened with toluol.
o. Place restraining strap in position and secure
with screws, washers and sleeves.
15-32A. WINDSHIELD ANTI-ICE PANEL (REMOVABLE.) (See figure 15-7B.)
15-32B. DESCRIPTION. Thru 1977 models, thepanel
is constructed of two sheets of plate glass covering a
layer of vinyl. Imbedded in the vinyl is a fine resis-
tance wire which provides the heat for windshield deicing. The lower edge of the panel is mounted on the
deck skin just forward of the windshield. The upper
end of the panel is supported by a rubber bumper
which holds the panel off the windshield. The lower
mounting bracket is hinged for easy cleaning between
the panel and windshield. The hinge pins are spring
loaded so the panel may be easily removed. Power
to the windshield panel is provided through a plug located in a housing assembly just left of the lower support bracket. A drain tube is provided for the housing assembly also a plug button is provided, which is
painted the same color as the deck skin, to plug connector hole in the deck skin when the anti-ice assembly is removed. A circuit breaker switch located on
the instrument panel is a off-on switch and a circuit
breaker to protect the system. Beginning with 1978
models the panel extends the full height of the windshield. The upper and lower ends of the panel are
held in place by retainers and screws. The system
is controlled by a rocker switch on the instrument
panel which connects power to the controller from a
15 amp circuit breaker on the bus bar. The controller is mounted on the glove box. Power is also fed
from the circuit breaker to a normally open relay,
also mounted on the glove box. The controller senses
the temperature of the panel and closes which feeds
power to the relay coil, closing the relay and power
is fed to the panel. When not in use the panel may be
removed and stowed in the aircraft.
15-32C. REMOVABLE AND INSTALLATION. (See
figure 15-7B.) Beginning with 1978 Models, when
the panel is removed and stowed, replace the AN
509-8R16 screws with AN509-8R12 screws. Also,
replace cover (8) with cover (11) (Figure 15-7B,
sheet 3.)
15-32D. HEATED WINDSHIELD PANEL (FIXED.)
15-32E. DESCRIPTION. An optional heated panel
is provided to prevent ice formation on the windshield.
The system consists of an electrically heated panel
attached to the windshield, a controller and a relay
mounted on the glove box. The system is controlled
by a rocker type switch on the pilot's switch panel.
A circuit breaker on the circuit breaker panel protects the system.
15-32F. REMOVAL AND INSTALLATION. (See figure 15-7B, sheet 3.)
a. Panel Removal.
1. Ensure aircraft electrical power is "OFF".
2. Disconnect housing plug and cap, located
forward of instrument panel on the left hand side.
3. Remove screws securing cover and gasket
to deck skin, then pull housing plug up through skin.
4. Remove screws from retainers at top and
bottom of heated panel.
5. Remove heated panel, retainers and shims
at top and bottom of panel.
6. Remove any sealer that may have parted
sticking to the windshield. A sharpened (Wood)
spatula may be used, exercising care.
Revision 1
15-40A
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
12. Place the drilling shield between heated panel
and windshield retainer and drill (. 172) holes at the
marked locations.
13. Place the upper spacer in position between
heated panel and windshield and temporarily secure
using three screws.
14. Check the temporary installation to ensure
that heated panel is in proper relation to the windshield. Check to see if panel seal is in contact with
windshield.
15. Remove the masking tape applied to windshield for locating heated panel. Apply new strips
of masking tape on each side of the panel with edge
Do Not use any tool, abrasive or cleaner
which may damage the windshield.
b.
Panel Installation.
1. Apply a strip of masking tape on the LH
windshield, from top to bottom with outboard edge
of tape located 6. 60 inches to the left and parallel
with the windshield conterline, as viewed looking
forward.
2. Apply a strip of masking tape at the bottom
of heated panel location with edge running parallel
with, and . 55 inch below the center of the three open
fastener locations. However, this dimension may
vary as lower edge of heated panel may be trimmed
to match aircraft contours. A minimum of .35 inch
edge margin must be maintained.
3. Locate heated panel with lower end and inboard side against edge of masking tape. Using a
hole finder, locate and mark the three hole locations
at the lower end of the panel
4. Drill three .172 holes on the lower end of the
panel where marked.
5. Place lower spacer in position and temporarily secure the lower end of heated panel with
three screws.
6. Press the heated panel to the windshield contour working up from the bottom so that panel seal is
compressed against windshield, firmly tape heated
panel to the windshield.
NOTE
aligned with and against outer lip of seal to facilitate
final installation. Also apply strips of tape at upper
and lower edge of heated panel.
16. Remove heated panel and deburr all parts.
17. Remove protective cover from the heated
panel. Do not remove masking tape aligning guides.
Clean thoroughly with a soft cloth or sponge. Wash
with a mild soap and water, a 50/50 solution of
isopropanol and water, or aliphatic naptha type 2.
Do not use any abrasive materials, strong acid or
base, methanol or methyl-ethyl-ketone. After cleaning, rinse thoroughly and dry.
18. After cleaning, plastic surfaces may be polished by applying a thin coat of hard polishing wax.
Rub lightly with a soft cloth using a circular motion.
19. Apply a bead of RTV108 sealer to the groove
of heated paneL
NOTE
Do not allow the RTV108 sealer to be pressed
out of the seal upon installation. If this happens,
remove heated panel, wipe the sealer off the
windshield and the seal on the heated panel
with isopropyl alcohol. Reapply RTV108 sealer
in groove, correcting the amount of bead, and
The inner and outer lip of the heated panel
seal should be in positive contact with the
surface of the windshield over the full periphery of the panel. It is permissible to
vary thickness of the spacers to facilitate
proper sealing.
reinstall the heated panel.
7. Using a hole finder, mark the center hole
location at the upper end of panel.
NOTE
Before drilling three .172 diameter holes
in the upper end of panel, place a metal
shield between the panel and windshield of
aircraft to protect the windshield from
damage.
8. Locate and drill one (.172) diameter hole
0.10 inch down from the mark on the heated panel.
9. Remove drilling shield.
10. Use an ice pick to aling hole in heated panel
with open hole in windshield retainer, and pull panel
up to align holes.
NOTE
20. Install heated panel on windshield exercising
care to prevent smearing of sealer.
21. Ensure proper location of spacers at upper
and lower ends of heated panel. (See note after step
5).
22.
Install screws at top and bottom of heated
panel.
23. Route heated panel electrical leads through
the deck skin and gaskets then connect.
24. Install cover and apply a strip of tape around
opening to keep sealer off of deck skin. Apply RTV108
sealer, potting wire bundle in cover.
NOTE
Allow 24 hours for full cure of RTV108 sealer.
Take precaution to prevent damage to windshield and/or doubler nutplates when tightening heated panel on windshield.
25. Remove all tape around heated panel and lead
cover.
26. Operational check the heated panel as follows:
a. Turn windshield de-ice switch momentarily ON,
check ammeter for discharge.
11. Using a hole finder, mark the remaining two
holes at the upper end of the panel
15-32G. TRAPPED MOISTURE. To eliminate moisture trapped between the heated windshield panel and
Revision 1
15-40E
MODEL 210 & T210 SERIES SERVICE MANUAL
the windshield, proceed as follows:
1. Loss of outlet set pressure.
a. Fabricate two probes from .125 diameter tube
2. Loss of oxygen flow through the regulaapproximately three inches long. Cut one end of
tor which will result in inadequate oxygen being fed
tubes off at approximately a 300 or less angle. File
through the aircraft system.
to a sharp edge.
3. Internal leakage of oxygen through regub. Insert one tube through the upper outboard corlator.
ner of the heated panel and the other through the
lower inboard corner. Move lower tube to the outOpening of the control lever with the outlet ports
board corner as required to release all trapped water.
open to atmosphere, results in an "overshoot" of
Insert tubes through the rubber seal.
the regulator metering device due to the extreme
c. Connect upper tube to a source of low pressure
flow
demand through the regulator. After overshootdry air, or bottled nitrogen. Flow air between the
ing, the metering poppet device goes into oscillation,
heated panel and windshield until all visible moisture
creating serious damage to the poppet seat and diais gone. Activate heated panel for short periods to
phragm metering probe. This condition can occur
accelerate removal of moisture.
even by turning the control lever on and then turning
d. Apply soap and water mixture to edges of the
it quickly off.
heated panel. Restrict exit air, noting and marking leakage from under panel. Do not overpresA potential hazard exists to aircraft in the field
sure; use no more than 2.0 psi.
where inexperienced personnel might remove the
e. Clean windshield and edge of heated panel with
cylinder and regulator assembly from the aircraft
mild soap and water and a 50/50 solution of isopropyl
and for some reason attempt to turn the regulator
alcohol and water. Wipe dry and apply masking tape
to the "ON" position with the outlet ports open. Unalong leak area approximately . 06-inch from seal.
fortunately, after the units have been improperly
Lift edge of seal and insert RTV. Fill gap at upper
operated as noted, there is no outward appearance
and lower ends of heated panel between panel seat
indicating that damage has occurred.
and windshield retainer with RTVif leak is in this
area. Remove tubes from windshield; fill holes with
Testing these regulators should be accomplished only
RTV and remove masking tape. Use clear RTVafter installation in the aircraft, with the "downstream" low pressure line attached.
108 only.
15-33.
OXYGEN SYSTEM.
(See figure 15-8.)
WARNING
Under No circumstances, turn the ON-OFF
control to the "ON" position with the outlet
(low pressure) ports open to atmosphere.
This action will induce serious damage to
the regulator, with the following results:
15-40F
Revision 2
15-34. DESCRIPTION. The system is comprised of
four oxygen cylinders, mounted in the cabin top area,
_in front of and behind the main carry-thru spar. Of
assembly. Remaining components of the system include a filler valve, located in the lower inboard surface of the right wing, cabin outlets, mask assemblies, and a pressure gage at the pilot's position.
The pilot's supply line is designed to receive a greater flow of oxygen than the passengers. The pilot's
MODEL 210 & T210 SERIES SERVICE MANUAL
mask is equipped with a microphone, keyed by a
switch button on the pilot's control wheel. An ONOFF control is provided at the pilot's position.
NOTE
Most air compressors are oil lubricated,
and a minute amount of oil may be
WARNING
WARNING
Oil, grease or other lubricants in contact
with
oxygen, create
create aa seriwith high-pressure
high-pressure oxygen,
serious fire hazard and such contact should be
avoided. Do not permit smoking or open
flame in or near aircraft while work is performed on oxygen systems.
1535. MAINTENANCE PRECAUTIONS.
a. Working area, tools and hands must be clean.
b. Keep oil, grease, water, dirt, dust and all
other foreign matter from system.
c. Keep all lines dry and capped until installed.
d. Use only MIL-T-5542 thread compound or teflon
lubricating tape on threads of oxygen valves, tubing
connectors, fittings, parts of assemblies which might
under any conditions, come in contact with oxygen.
The thread compound must be applied sparingly and
carried by the airstream. If only an oil
lubricated air compressor is available,
drying must be accomplished by heating
drying
must be accomplished
by heating
at
of
(121º
at aa temperature
temperature
of 250º
250° to
to 300ºf
300ºF
(121º to
to
149°C) for a suitable period.
2. Flush with naphtha, conforming to Specification TT-N-95 (aliphatic naphtha). Blow clean and
dry off all solvents with clean, fluid,
oil-free, filtered air. Flush with anti-icing fluid conforming to
Specification TT-T-735 or anhydrous ethyl alcohol.
Rinse thoroughly with fresh water. Dry thoroughly
with a stream of clean, dry, oil-free, filtered air.
3. Flush with hot inhibited alkalinecleaner until free from oil and grease.
inse with fresh water
and dry with clean, dry, filtered air.
NOTE
carefully to only the first three threads of the male
fitting. No compound shall be used on aluminum
Cap lines at both ends immediately after
to prevent contaimination.
flared fittings or on the coupling sleeves or on the
used in accordance with the instructions listed following this step, Extreme care must beexercisassembly,
to prevent the contamination of the thread compound
or teflon tape with oil, grease or other lubricant.
1. Place tape on threads close to end of fitting.
Wrap clockwise on RH threads, counterclockwise on LB threads.
2. Apply enough tension while winding so tape
forms into thread grooves.
3. After wrap is complete, maintain tension
and tear tape by pulling apart in direction
it was applied. Resulting ragged end is the
key to the tape staying in place. (If sheared
or cut, tape may unwind.)
4. Press tape well into threads.
5. Make connections.
e. Fabrication of oxygen pressure lines is not
recommended. Lines should be replaced by part
numbers called out in the aircraft Parts Catalog.
f. Lines and fittings must be clean and dry. One
of the following methods may be used.
1. Clean by degreasing with stabilized trichlorethylene, conforming to Federal Specifications
O-T-634 or MIL-T-27602. These items can be
tained from American Mineral Spirits of Houston,facility.
Texas.
15-36.
REPLACEMENT OF COMPONENTS. Rea
and install accomplished
figure 15-8 as a guide.
CAUTION
Oxygen cylinders and regulators are
furnished as assemblies by Cessna Parts
Distribution (CPD 2). Attempting to
remove, repair, and reinstall oxygen
regulators in the field provides
opportunity for contaminants to enter
the system. Faulty regulators or
regulators otherwise in need of
disassembly should be exchanged for
replacement oxygen bottle and regulator
assemblies through CPD 2. Regulator
and cylinder assembly shall be
disassembled, repaired, inspected,
cleaned, hydrostatically tested,
reassembled, and serviced by
manufacturer or other FAA-approved
other
Revision 3
15-41
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Oxygen cylinder and regulator assemblies
may not always be installed in the field
exactly as illustrated in figure 15-8, which
shows factory installation. Important
points to remember are as follows.
a. Before removing cylinder, release low-pressure line by opening cabin outlets. Disconnect pushpull control cable, filler line, pressure gage line
and outlet line from regular. CAP ALL LINES
IMMEDIATELY.
b. If it is necessary to replace filler valve O-rings,
remove parts necessary for access to filler valve.
Remove line from quick-disconnect valve at the
regulator, then disconnect chain, but do not remove
cap from filler valve. Remove screws securing
valve and disconnect pressure line. Referring to
applicable figure, cap pressure line and seat. Diassemble valve, replace O-rings and reassemble
valve. Install filler valve by reversing procedures
outlined in this tep.
c. To remove entire oxygen system, headliner
must be lowred and soundproofing removed to expose lines. Refer to Section 3 for headliner remoral.
15-37. OXYGEN CYLINDER GENERAL INFORMATION. The following nformation is permanently
steel stamped on the shoulder, top head or neck of
each oxygen cylinder:
a. Cylinder specification, followed by service
pressure (e.g. '1CC-3AA1800 and CC-3HT1850"
for standard and light weight cylinders respectively).
NOTE
Effective 1 January 1970, all newly- manufactured cylinders are stamped "DOT" (De-
partment of Transportation), rather than
'CC" (Interstate Commerce Commission).
An example of the new designation would be:
"DOT-3HT1850".
b. Cylinder serial number is stamped below or
directly following cylinder specification. The symbol of the purchaser, user or maker, if registered
with the Bureau of Explosives, may be located directly below or following the serial number. The
cylinder serial number may be stamped in an alternate location on the cylinder top head.
c. Inspector's official mark near serial number.
d. Date of manufacture: This is the date of the
first hydrostatic test (such as 4-69 for April 1969).
The dash between the month and the year figures
may be replaced with the mark of the testing or inspection agency (e.g. 4L69).
e. Hydrostatic test date: The dates of subsequent
hydrostatic tests shall be steel stamped (month and
year) directly below the original manufacture date.
The dash between the month and year figures can be
replaced with the mark of the testing agency.
f. A Cessna identification placard is located near
the center of the cylinder body.
g. Halogen test stamp: "Halogen Tested", date of
test (month, day and year) and inspector's mark
SHOP NOTES:
15-43
MODEL 210 & T210 SERIES SERVICE MANUAL
1
Detail
4. Cap
Filler
5.
andValve
Chain
A
14.
15.
Bracket
Cover
Detail C
Used With
Air Conditioning
Used
WithAirConditoning
11
1. Filler Line
2. O-Ring
11. Casing
3.
4.
5.
6.
7.
Bracket
Filler Valve
Cap and Chain
Cover
Tee
13.
14.
15.
16.
17.
Cover Assembly
Regulator
Wire
18. Spacer
19. Control Lever
20. Knob Assembly
8.
9.
10.
12. Adapter
Cylinder
Bracket
Cover
Valve Assembly
Pressure Gage
Figure 15-10.
14
12
14
Emergency Oxygen System
13
(Sheet 1 of 2)
15-45
MODEL 210 & T210 SERIES SERVICE MANUAL
1
18
-101
i
18
'19
18
14
Detail F
Detail
E
Used with Air Conditioning
14
17
'
Detail
Figure 15-10. Emergency Oxygen System (Sheet 2 of 2)
15-46
O
MODEL 210 & T210 SERIES SERVICE MANUAL
appears directly underneath the Cessna identification
placard.
tive January 17. 1978, does not apply to SP5957 cylinders, even if these cylinders are marked as 3HT
cylinders. Such cylinders can be identified by the
marking "SP5957", which will appear on the shoulder
of the cylinder. Any cylinder so marked, regardless
of any other markings that may also appear. is not a
DOT 3HT cylinder. and the service life extension from
15 years to 24 years does not apply.
15-38. OXYGEN CYLINDER SERVICE REQUIREMENTS.
a. Hydrostatic test requirements:
1. Standard weight (ICC or DOT-3AA1800)
cylinders must be hydrostatically tested to 5/3 their
working pressure every five years commencing with
15-40. OXYGEN SYSTEM COMPONENT SERVICE
the date of the last hydrostatic test.
REQUIREMENTS.
2. Light weight (ICC or DOT-3HT1850) cylina. PRESSURE REGULATOR. The regulator shall
ders must be hydrostatically tested to 5/3 their
be removed and overhauled by manufacturer or an
working pressure every three years commencing
FAA approved facility during hydrostatic testing.
with the date of the last hydrostatic test.
b. Service life requirements:
1. Standard weight (ICC or DOT-3AA1800)
CAUTION
cylinders have no age life limitations and may continue to be used until they fail hydrostatic test.
Oxygen cylinders and regulators are
2. Light weight (ICC orDOT-3HT 1850) cylinfurnished as assemblies by Cessna Parts
ders must be retired from service after 24 years or
Distribution (CPD 2). Attempting to
4,380 filling cycles after date of manufacture, which-remove,
repair, and reinstall oxygen
ever occurs first. If a cylinder is recharged more
regulators in the field provides
than an average of once every other day, an accurate
opportunity for contaminants to enter
record of the number of rechargings must be mainthe system. Faulty regulators or
regulators otherwise need
tatned. Refer to paragraph 15-39 for determining
regulators otherwise in need of
service life of DOT-3HT1850 cylinders.
disassembly should be exchanged for
replacement oxygen bottle and regulator
assemblies through CPD 2. Regulator
NOTE
and cylinder assembly shall be
disassembled, repaired, inspected,
These test periods and life limitations
cleaned, hydrostatically tested,
are established by the Department of
reassembled, and serviced by
Transportation Code of Federal Regumanufacturer or other FAA-approved
lations; Title 49, Chapter 1, Para.
facility.
73.34.
15-39. OXYGEN CYLINDER INSPECTION REQUIREMENTS.
a. Inspect the entire exterior surface of the cylinder for indication of abuse, dents, bulges and strap
chafing.
b. Examine the neck of cylinder for cracks, distortion or damaged threads.
c. Check the cylinders to determine if markings
are legible.
d. Check date of last hydrostatic test. If the periodic retest date is past, do not return the cylinder
to service until the test has been accomplished.
e. Inspect the cylinder mounting bracket, bracket
hold-down bolts and cylinder holding straps for
cracks, deformation, cleanliness, and security of
attachment.
f. In the immediate area where the cylinder is
stored or secured, check for evidence of any types
of interference, chafing, deformation or deterioration.
g. A cylinder manufactured prior to January 17,
1978, and not yet marked with a rejection elastic expansion (REE), must be marked with that REE in cubic centimeters near the marked original elastic expansion prior to the next retest date. The REE for a
cylinder is 1. 05 times its original elastic expansion.
h. Some cylinders manufactured to DOT special permit 5957 in the past. were incorrectly marked with
"DOT 3HT" in addition to "SP5957". Cylinders made
under SP5957 are not DOT3HT cylinders, and the service life extension from 15 years to 24 years, effec-
b. FILLER VALVE. The valve should be disassembled, inspected and the O-rings replaced, regardless
of condition, every 3 years or 3000 flight hours,
whichever occurs first.
c. QUICK-RELEASE COUPLING. The coupling
shall be functionally tested every two years and overhauled every five years or at time of hydrostatic
test.
d. PRESSURE GAGE. The gage shall be replaced
when found to be faulty. No re-conditioning or overhaul of the gage is authorized.
e. INDIVIDUAL OUTLETS. The outlets shall be
disassembled and inspected and the O-rings replaced,
regardless of condition, every 3 years or 3000 flight
hours, whichever occurs first.
15-41. OXYGEN SYSTEM COMPONENT INSPECTION REQUIREMENTS.
a. Examine all parts for cracks, nicks, damaged
threads or other apparent damage.
b. Actuate regulator controls and valve to check
for ease of operation.
c. Determine if the gage is functioning properly by
observing the pressure buildup and the return to zero
when the system oxygen is bled off.
d. Replace any oxygen line that is chafed, rusted,
corroded, dented, cracked or kinked.
e. Check fittings for corrosion around the threaded area where lines are joined together. Pressurize the system and check for leaks.
Revision 3
15-47
I
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
Each interconnected series of oxygen cylinders is equipped with a single gage. The trailer type
cascade may also be equipped with a nitrogen cylinder (shown reversed) for filling landing gear
struts, accumulators, etc. Cylinders are not available for direct purchase, but are usually
leased and refilled by a local compressed gas supplier.
Service Kit SK310-32 (available from Cessna Parts Distribution [CPD 2] through Cessna Service
Stations) contains an adapter, pressure gage, and hose, and lines and fittings for equipping two
oxygen cylinders to service oxygen systems. As noted in the Service Kit, a tee (Part No. 11844) and a
pigtail (Part No. 1243-2) should be ordered for each additional cylinder to be used in the cascade of
cylinders. Be sure to ground the airplane and ground servicing equipment before use.
OXYGEN CYLINDER
PRESSURE GAGE
CYLINDER
OXYGEN PURIFIERW/REPLACEABLE
CARTRIDGE
Figure 15-11.
Typical Portable Oxygen Cascades
15-42. MASKS AND HOSE.
a. Check oxygen masks for fabric cracks and rough
face seals. If the mask is a full-faced model, inspect glass or plastic for cleanliness and state of
repair.
b. Flex the mask hose gently over its entirety and
check for evidence of deterioration or dirt.
c. Examine mask and hose storage compartment
for cleanliness and general condition.
15-43. MAINTENANCE AND CLEANING.
a. Clean and disinfect mask assemblies after use,
as appropriate.
NOTE
Use care to avoid damaging microphone
assembly while cleaning and sterilizing.
b. Wash mask with a mild soap solution and rinse
it with clear water.
c. To sterilize, swab mask thoroughly with a
gauze or sponge soaked in a water/merthiolate solution. This solution should contain 1/5 teaspoon of
merthiolate per one quart of water. Wipe the mask
with a clean cloth and let air dry.
d. Observe that each mask breathing tube end is
free of nicks and that the tube end will slip into the
cabin oxygen receptacle with ease and will not leak.
e. If a mask assembly is defective (leaks, does not
allow breathing or contains a defective microphone)
it is advisable to return the mask assembly to the
manufacturer or a repair station.
15-48
Revision 3
Replace hose if it shows-evidence of deteriof.
ration.
g. Hose may be cleaned in the same manner as the
mask.
15-44. SYSTEM PURGING. Whenever components
have been removed and reinstalled or replaced, it is
advisable to purge the system. Charge oxygen system in accordance with procedures outlined in paragraph 15-47. Plug masks into all outlets and turn
the pilot's control to ON position and purge system
by allowing oxygen to flow for at least 10 minutes.
Smell oxygen flowing from outlets and continue to
purge until system is odorless. Refill cylinders as
required during and after purging.
15-45. FUNCTIONAL TESTING. Whenever the regulator and cylinder assembly has been replaced or
overhauled, perform the following flow and internal
leakage tests to check that the system functions properly.
a. Fully charge oxygen system in accordance with
procedures outlined in paragraph 15-47.
b. Disconnect line and fitting assembly from pilot's mask and line assembly. Insert outlet end of
line and fitting assembly into cabin outlet and attach
opposite end of line to a pressure gage (gage should
be calibrated in one-pound increments from 0 to 100
PSI). Place control lever in ON position. Gage
pressure should read 70*10 PSI.
c. Insert mask and line assemblies into all remaining cabin outlets. With oxygen flowing from all
outlets, test gage pressure should still be 70*10 PSI.
MODEL 210 & T210 SERIES SERVICE MANUAL
d. Place oxygen control lever in OFF position and
allow test gage pressure to fall to 0 PSL Remove
all adapter assemblies except the one with the pressure gage. The pressure must not rise above 0 PSI
when observed for one minute. Remove pressure
gage and adapter from oxygen outlet.
CAUTION
A cylinder which is completely empty may
well be contaminated. The regulator and
cylinder assembly must then be disassembled, inspected and cleaned by an FAA
approved facility, before filling. Contamination, as used here, means dirt, dust
or any other foreign material, as well as
ordinary air in large quantities. If a gage
line or filler line is disconnected and the
fittings capped immediately, the cylinder
will not become contaminated unless temperature variation has created a suction
within the cylinder. Ordinary air contains
water vapor which could condense and
freeze. Since there are very small orifices
in the system, it is very important that
this condition not be allowed to occur.
NOTE
If pressures specified in the foregoing procedures are not obtained, the oxygen regulator is not operating properly. Remove
and replace cylinder-regulator assembly
with another unit and repeat test procedure.
e. Connect mask and line assemblies to each cabin
outlet and check each mask for proper operation.
f. Check pilot's mask microphone and control
wheel switch for proper operation. After checking,
return all masks to mask case.
g. Recharge oxygen system in accordance with
procedures outlined in paragraph 15-47.
15-46. SYSTEM LEAK TEST. When oxygen is being
lost from a system through leakage, a sequence of
steps may be necessary to locate the opening. Leakage may often be detected by listening for the distinct hissing of escaping gas. If this check proves
negative, it will be necessary to soap-test all lines
and connections with a castile soap and water solution or specially compounded leak-test material.
Make the solution thick enough to adhere to the contours of the fittings. At the completion of the leakage test, remove all traces of the leak detector or
soap and water solution.
CAUTION
1to
Do not attempt to tighten any connections
while the system is charged.
15-47.
15-47
CHARGING
SYSTEM
SYSTEM CHARGING.
WARNING
BE SURE TO GROUND AIRCRAFT AND
GROUND SERVICING EQUIPMENT BEFORE CHARGING OXYGEN SYSTEM.
a. Do not attempt to charge oxygen cylinders if
servicing equipment fittings or filler valve are
corroded or contaminated. If in doubt, clean with
stabilized trichlorethylene and let air dry. Do not
allow solvent to enter any internal parts.
b. If cylinder is completely empty, do not charge,
as the cylinder must then be removed, inspected
and cleaned.
c. Connect cylinder valve outlet or outside filler
valve to manifold or portable oxygen cascade.
d. Slowly open valve on cascade cylinder or manifold with lowest pressure, as noted on pressure gage,
allow pressure to equalize, then close cascade cylinder
valve.
e. Repeat this procedure, using a progressively
higher pressure cascade cylinder, until system has
been charged to the pressure indicated in the chart
immediately following step "f" of this paragraph.
f. Ambient temperature listed in the chart is the
air temperature in the area where the system is to
be chrged Filling pressure refers to the pressure to which aircraft cylinders should be filled
sure to which aircraft cylinders should be filled.
This table gives approximations only and assumes
a rise in temperature of approximately 25°F. due
heat of compression. This table also assumes
the aircraft cylinders will be filled as quickly as possible and that they will only be cooled by ambient
air; no water bath or other means of cooling be used.
Example: If ambient temperature is 70°F., fill
aircraft cylinders to approximately 1, 975 psi or as
close to this pressure as the gage may read. Upon
cooling, cylinders should have approximately 1, 850
pressure.
psi
TABLE
OF FILLING PRESSURES
Ambient
Temp.
ºF
10
20
30
40
50
60
Filling
Press.
psig
165
1700
1725
1775
1825
1875
1925
Ambient
Temp.
°F
0
Revision 3
0
80
90
100
110
120
130
Filling
Press.
psig
1975
1975
2000
2050
2100
2150
2200
2250
15-49/(15-50 blank)l
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 16
INSTRUMENTS AND INSTRUMENT SYSTEMS
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
INSTRUMENTS AND
INSTRUMENTSYSTEMS ..........
General ......................
Instrument Panel ..............
Description .................
Removal and Installation .....
Shock-Mounts ............
Instruments ..............
Removal ...............
Installation ............
Pitot and Static Systems ........
Description .................
Maintenance ................
Static Pressure System
Inspection and Leakage Test .
Pitot System Inspection and
Leakage Test ..............
Blowing Out Lines ...........
Removal and Installation of
Components ...............
*
I
Trouble Shooting - Pitot-Static
System ....................
I
True Airspeed Indicator ......
Trouble Shooting ..........
Trouble Shooting- Altimeter ..
Trouble Shooting - Vertical
Speed Indicator ............
Trouble Shooting - Pitot
..............
Tube Heater .
Vacuum System ................
................
Description .
Trouble Shooting - Vacuum
.............
System .
Trouble Shooting - Gyros .....
Trouble Shooting -Vacuum
Pump .....................
Maintenance Practices .......
Removal of Vacuum Pump ....
Mounting Pad Inspection .....
Installation of Vacuum Pump .
Cleaning ....................
Low-Vacuum Warning Light ..
Vacuum Relief Valve
Adjustment ................
Engine Indicators ..............
Tachometer .................
Description ...............
Manifold Pressure/Fuel Flow
Indicator ..................
Description ...............
Trouble Shooting - Manifold
Pressure Indicator .......
Trouble Shooting - Fuel Flow
Indicator ................
2K1/161
2K1/161
2K7/16-3
2K7/16-3
2K7/16-3
2K7/16-3
2K7/16-3
2K7/16-3
2K7/16.3
2K7/16-3
2K716-3
2K7/16-3
2K10/16-6
2K10/16-6
2K10/16.6
2K11/16-7
2K11/16-7
Removal and Installation .
Transmitter Calibration ...
2K16/16-12
2K16/16-12
Stewart Warner Gage
Transmitter Calibration
Rochester Gage
Transmitter .
Hourmeter ..................
Description ...............
Economy Mixture Indicator ...
Description ...............
Trouble Shooting ..........
Calibration ...............
Removal and Installation
Magnetic Compass ...........
Description ...............
Stall Warning Horn ..........
Description ...............
2K16/16-12
Turn Coordinator ............
2K11/16-7
2K11/16-7
2K13/16-9
2K13/16-9
2K14/16-10
2K14/16-10
2K14/16-10
2K14/16-10
2K15/16-11
2K16/16-12
2K16/16-12
2K19/16-15
2K19/16-15
2K20/16-16
2K20/16-16
2K20/16-16
2K20/16-16
2K20/16-16
2K20/16-16
2K21/16-17
2K22/16-18
18-1. INSTRUMENTS AND INSTRUMENT SYSTEMS.
16-2. GENERAL.
Cylinder Head Temperature
age ......................
Description ...............
Trouble Shooting ..........
Oil Pressure Gage ............
Description ...............
Trouble Shooting ..........
Oil Temperature Gage
Description ...............
Fuel Quantity Indicating
System (Thru 21062273) ....
Indicators ................
Sending Units .............
Contro Monitor ...........
Removal and Installation ..
Calibration ...............
Trouble Shooting ..........
Fuel Quantity Indicating System
(Beginning with 21062274) ..
Description ...............
Trouble Shooting ..........
This section describes typical in-
strument installations and their respective operating
systems. Emphasis is placed on trouble shooting and
corrective measures only. It does NOT deal with specific instrument repairs since this usually requires
special equipment and data and should be handled by
instrument specialists. Federal Aviaton Regulations
2K22/16-18
2K22/16-18
2K22/16-18
2K23/16-19
2K23/16-19
2K23/16-19
2K24/16-20
2K24/1620
2K24/16-20
2K24/1620
2K24/16-20
2K24/16-20
2K24/16-20
2K24/16-20
2K24/16-20
2L1/16-21
21/16-21
2L116-21
2L2/16-22
2L2/16-22
2L2/16-22
2L216-22
2L5/16-23
2L5/16-23
2L5/16-23
2L5/16-23
2L5/16-23
2L5/16-23
2L6/16-24
2L6/16-24
2L6/16-24
2L6/16-24
2L6/16-24
2L6/16-24
Description ...............
Trouble Shooting ..........
Turn and Slip Indicator .......
Description ...............
Trouble Shooting ..........
Electric Clock ................
Description ...............
Fuel Computer/Digital Clock .
Description ...............
Fuel Computer Operation
Digital Clock Operation .
Trouble Shooting ..........
Fuel Flow Transducer ......
Installation ............
Removal and Replacement
Calibration ............
2L6/16-24
2L6/16-24
2L8/16-26
2L8/16-26
2L8/16-26
2L9/16-27
2L9/16-27
2L9/16-27
2L9/16-27
2L9/16-27
2L11/16-29
2L11/16-29
2L12/16-30
2L12/16-30
2L13/16-31
2L14/16-32
require malfunctioning instruments be sent to an approved instrument overhaul and repair station or returned to manufacturer for servicing.
Our concern
here is with preventive maintenance on various instrument systems and correction of system faults which
result in instrument malfunctions. The descriptive
material, maintenance and trouble shooting information in this section is intended to help the mechanic
Revision 3
16-1
MODEL 210 & T210 SERIES SERVICE MANUAL
6
.7E10
4987
16
6
a
14¶
s
168
12)~ \~
18
'X
B
Detil A
B
7. Headg and Ventiatin
1.
2.
3.
4.
5.
6.
Marker Beacon Controls
Shock-Mounted Panel
Removeable Flight Instrument Panel
Radio and Switch Panel
Fuel and Engine Instruments
Protection Pad
8.
9.
10.
11.
12.
13.
Controls
Wing Flap Controls
Engine Controls
Circuit Breaker Panel
Ignition and Switch Panel
Nut
Lock Washer
Figure 16-1. Ins
14.
15.
16.
17.
18.
19.
Shockc-Mount
Ground Strap
Decorative Cover
Stationary Panel
Stud
Threaded Button
ent Panel (Typical)
16-2
14ry
\
a
^^*
<
s
-
MODEL 210 & T210 SERIES SERVICE MANUAL
determine malfunctions and correct them, up to the
defective instrument itself, at which point an instrument technician should be called in. Some instruments, such as fuel quantity and oil pressure gages,
are so simple and inexpensive, repairs usually will
be more costly than a new instrument. On the other
hand, aneroid and gyro instruments usually are well
worth repairing. The words "replace instrument"
in the text, therefore, should be taken only in the
sense of physical replacement in the aircraft. Whether replacement is to be with a new instrument, an
exchange one, or the original instrument is to be repaired must be decided on basis of individual circumstances.
16-3.
INSTRUMENT PANEL.
(Refer to figure 16-1.)
16-4. DESCRIPTION. The instrument panel assembly consists of a stationary panel, a removable flight
instrument panel and a shock-mounted panel. The
stationary panel, containing fuel and engine instruments is secured to the engine mount stringers and a
forward fuselage bulkhead. The removable panel,
containing flight instruments such as airspeed, vertical speed and altimeter is secured to the stationary
panel with screws. The shock-mounted panel, containing major flight instruments such as the horizontal and directional gyros is secured to the removable
panel with rubber shock- mounted assemblies. Most
of the instruments are screw mounted on the panel
*~
FLIGHT INSTRUMENT PANEL.
1. Unscrew threaded buttons holding decorative
cover.
2. Pull decorative cover back and disconnect
post light wires, if installed, and remove decorative
cover.
3. Tag and disconnect plumbing and wiring.
4. Remove screws securing flight instrument
panel to stationary panel and pull panel straight back.
5. Reverse preceding steps for reinstallation.
b. SHOCK-MOUNTED PANEL.
-NOTE
~
INSTRUMENTS.
(Refer to figure 16-1.)
16-8. REMOVAL. Most instruments are secured to
the panel with screws inserted through the panel face,
under the decorative cover. To remove an instrument,
remove decorative cover, disconnect wiring or plumbing to instrument, remove mounting screws and take
instrument out from behind, or in some'cases, from
front of panel. Instrument clusters are installed as
units and are secured by a screw at each end. A
cluster must be removed from the forward side of the
stationary panel to replace an individual gage. In all
cases when an instrument is removed, disconnected
lines or wires should be protected. Cap open lines
and cover pressure connections on instrument to prevent thread damage and entrance of foreign matter.
Wire terminals should be insulated or tied up to prevent accidental grounding or short-circuiting.
16-9. INSTALLATION. Generally, installation procedure is the reverse of removal procedure. Ensure
mounting screw nuts are tightened firmly, but do not
over-tighten, particularly on instruments having
plastic cases. The same rule applies to connecting
plumbing and wiring.
NOTE
All instruments (gages and indicators),
requiring a thread seal or lubriant.
shall be installed using teflon tape on
a.
0_
16-7.
from Cessna Parts Distribution (CPD)
through Cessna Service Stations
When replacing an electrical gage in an instrument
cluster assembly, avoid bending pointer or dial plate.
Distortion of dial or back plate could change the callbration of gages.
16-10. PITOT AND STATIC SYSTEMS. (Refer to
figure 16-2.)
16-11. DESCRIPTION. The pitot system conveys
ram air pressure to the airspeed indicator. The
static system vents vertical speed indicator, alti-
Due to the difficulty encountered when remov-
meter and airspeed indicator to atmospheric pressure through plastic tubing connected to the static
ing the shock-mounted panel with the gyros
installed, it is recommended that the directional gyro be disconnected and removed prior
to removal of the shock-mounted panel.
ports. A static line sump is installed at each source
button to collect condensation in the static system.
A pitot tube heater and stall warning heater may be
installed. The heating elements are controlled by a
1. Complete steps 1 and 2 above.
2. Tag and disconnect gyro plumbing.
3. Remove directional gyro mounting screws
andremove gyro from shock-mounted panel.
4. Remove shock-mount nuts and work shockmounted panel out from behind flight instrument panel.
The horizontal gyro may also be removed from shockmounted panel, if desired.
5. Reverse preceding steps for reinstallation.
switch at the instrument panel and powered by the
electrical system. A static pressure alternate
source valve may be installed in the static system
for use when the external static source is malfunctioning. This source is to be used only in emergenpressure
used as
as aa static
static source,
source cabin
causing
the
cies.cies. When
when used
is substituted for atmospheric pressure, causing the
instrument readings to vary from normal. This valve
also permits draining condensate from the static lines.
Refer to Pilot's Operating Handbook for flight operation
using alternate static source pressure.
16-12. MAINTENANCE. Proper maintenance of the
pitot and static system is essential for proper opera-
16-6. SHOCK-MOUNTS. Service life of shockmounted instruments is directly related to adequate
shock-mounting of the panel. If removal of shockmounted panel is necessary, check mounts for dedeterioration and replace as necessary.
Revision 3
16-3
MODEL 210 & T210 SERIES SERVICE MANUAL
Detail
A
B
5
SECOND ALTIMETER
Detail A
1. Altimeter
NOTE
Do not overtighten screws (14),
and do not lubricate any parts.
Use spacers (10) as required
for adequate friction on ring
assembly (12).
2.
3.
4.
5.
6.
7.
8.
9.
Bracket
10. Spacer
Figure 16-2.
16-4
Vertical Speed Indicator
Airspeed Indicator
Static Line (To Sump)
Pitot Line (To Pitot Tube)
Alternate Static Source
Valve
Line (To Airspeed)
Line (Alternate Air)
11.
Instrument Panel
12.
13.
14.
15.
16.
17.
18.
True Airspeed Ring
Retainer
Mounting Screw
Decorative Cover
Sump
Static Port
Heater Element (Heated
Pitot Only)
19.
Pitot Tube Mast Body
20. Connector
Pitot-Static System (Sheet 1 of 2)
MODEL 210 & T210 SERIES SERVICE MANUAL
3
10
TRUE AIRSPEED
INSTALLATION
11
MODEL 210 & T210 SERIES SERVICE MANUAL
tion of altimeter, vertical speed and airspeed indicators. Leaks, moisture and obstructions in the pitot
system will result in false airspeed indications, while
static system malfunctions will affect the 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.
. STATIC
PRESSURE SYSTEM INSPECTION
AND LEAKAGE TEST. The following procedure
AND LEAKAGE TEST. The following procedure
source opening. Figure 16-3 shows one method of
obtaining positive pressure.
CAUTION
Do not apply positive pressure with the airspeed indicator or vertical speed indicator
connected to the static pressure system.
l.
Slowly apply positive pressure until the altimeter
indicates a 500-foot decrease in altitude and maintain
this altimeter indication while checking for leaks..
of mild source
solution
a
and
water
outlines inspection and testing of the static pressure
LEAK-TEC or
system, assuming the altimeter has been tested and
leaks.
soap alocate
a solution of mild
watching
LEAK-TECfor
inspected in accordance with current Federal Aviation
tion Regulations.
Regulations.
a. Ensure that the static system is free from en-
bubbles to locate leaks.
watching for
m. Tighten leaking connections. Repair or replace
parts found defective.
parts found defective.
b.
deformations of
b. Ensure
Ensure that
that no
no alterations
alterations or
or deformations of
the airframe surface have been made which would
affect the relationship between air pressure in the.
static pressure system and true ambient static air
cators into the static pressure system andrepeat
leakage test per steps "c" thru "h".
PITOT
SYSTEM INSPECTION
trapped moisture and restrictions.
n. Reconnect the airspeed and vertical speed indi-
AND
LE
pressure for any flight configuration.
c.
Seal one static source port with pressure sensi-
tive tape.
This seal must be air tight.
piece of tape over the small hole in the lower aft end
of pitot tube, fasten a piece of rubber or plastic tub-
over pitot tube, close opposite end of tubing and
d. Close the static pressure alternate source valve, ing
slowly roll up tube until airspeed indicator registers
if installed.
Secure tube
in
minutes
few minutes
and after
after aa few
tube and
range. Secure
in cruise
cruise range.
e. Attach a source of suction to the remaining static
recheck airspeed indicator. Any leakage will have
pressure source opening. Figure 16-3 shows one
reduced the pressure in system, resulting in a lower
method of obtaining suction
suction,
airspeed
method
of obtaining
rebefore retubing before
Slowly unroll
unroll tubing
indication. Slowly
airspeed indication.
until the altimeter indicates
apply suction
f. Slowly
moving it, so pressure is reduced gradually. Othera 1000-foot increase in altitude.
wise instrument may be damaged. If test reveals a
CAUTION
When applying or releasing suction, do not16-15
do
suction not
When applying or releasing
the range of vertical speed indicaexceed
tor orindicator
airspeed
g. Cut off the suction source to maintain a "closed"
system for one minute. Leakage shall not exceed
100 feet of altitude loss as indicated on the altimeter.
h. If leakage rate is within tolerance, slowly release the suction source and remove the tape from
static port.
NOTE
leak in system, check all connections for tightness.
the pitot
LINES. Although
16-15. BLOWING
pitot
Although the
OUT LINES
BLOWING OUT
system is designed to drain down to the pitot tube
opening, condensation may collect at other points in
the system and produce a partial obstruction. To
clear the line, disconnect it at the airspeed ndicator.
Using low pressure air, blow from the indicator end
of line toward the pitot tube.
Like the pitot lines, static pressure lines must be
kept clear and connections tight. Static source sumps
collect moisture and keeps system clear. However,
when necessary, disconnect static line at first instrument to which it is connected, then blow the line clear
with low pressure air.
If leakage rate exceeds the maximum allow-
able, first tighten all connections, then repeat leakage test. If leakage rate still exceeds the maximum allowable, use still exallowable , use the following procedure.
i. Disconnect the static pressure lines from airspeed
indicator and vertical speed indicator. Use suitable
fittings to connect the lines together so the altimeter
is the only instrument still connected into the static
pressure system.
j. Repeat the leakage test to check whether the
static pressure system or the bypassed instruments
are the cause of leakage. If the instruments are at
fault, they must be repaired by an "appropriately
rated repair station" or replaced. If the static pressure system is at fault, use the following procedure
to locate leakage.
k. Attach a source of positive pressure to the static
16-6
CAUTION
Never blow through pitot or static lines toward
instruments. Insure that (avionics) altitude sensor line is disconnected from static lines before
blowing out lines, or damage to sensor may occur.
NOTE
On aircraft equipped with an alternate static
source, use the same procedure, opening
the alternate static source valve momentarily
to clear line, then close valve and clear the
remainder of system.
Check all static pressure line connections for tightness. If hose or hose connections are used, check
them for general condition and clamps for security.
_
MODEL 210 & T210 SERIES SERVICE MANUAL
Replace hose which have cracked, hardened or show
other signs of deterioration.
16-16. REMOVAL AND INSTALLATION OF COMPONENTS. (Refer to figure 16-2). To remove the
pitot mast, remove the four mounting screws on the
side of connector (19) and pull mast out of connector
far enough to disconnect pitot line (5). Electrical
connections to the heater assembly (if installed) may
be disconnected through the wing access opening just
inboard of mast. Pitot and static lines are removed
in the usual manner, after removing wing access
plates, lower wing fairing strip and upholstery as
16-17.
required. Installation of tubing will be simpler if a
guide wire 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 pitot 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 overtightening and distorting
fittings. If twisting of plastic tubing is encountered
when tightening fittings, VV-236 (USP Petrolatum),
may be applied sparingly between tubing and fittings.
TROUBLE SHOOTING--PITOT-STATIC SYSTEM.
TROUBLE
PROBABLE CAUSE
REMEDY
LOW OR SLUGGISH AIRSPEED
INDICATION. Normal altimeter
and vertical speed.
Pitot tube deformed, leak or
obstruction in pitot line.
Straighten tube, repair or replace
damaged line.
INCORRECT OR SLUGGISH
RESPONSE. Al three instruments.
Leaks or obstruction in static
line.
Repair or replace line.
Alternate static source valve
open.
Close for normal operation.
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 the instrument is
read as true airspeed on the adjustable ring. Refer
to figure 16-2 for removal and installation. Upon in16-19.
TROUBLE SHOOTING.
stallation, before tightening mounting screws (14),
calibrate the instrument as follows: Rotate ring (12)
until 105 knots on adjustable ring aligns with 105 knots
on indicator. Holding this setting, move retainer (13)
until 60°F aligns with zero pressure altitude, then
tighten mounting screws (14) and replace decorative
cover (15).
NOTE
Refer to paragraph 16-15 before blowing out pitot or
static lines.
TROUBLE
HAND FAILS TO RESPOND.
INCORRECT INDICATION OR
HAND OSCILLATES.
PROBABLE CAUSE
REMEDY
Pitot pressure connection
not properly connected to
pressure line from pitot tube.
Repair or replace damaged line,
tighten connections.
Pitot or static lines clogged.
Blow out lines.
Leak in pitot or static lines.
Repair or replace damaged
lines, tighten connections.
Defective mechanism.
Replace instrument.
Leaking diaphragm.
Replace instrument.
Alternate static source valve
open.
Close for normal operation.
16-7
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
16-19.
TROUBLE SHOOTING (Cont).
TROUBLE
HAND VIBRATES.
16-20.
PROBABLE CAUSE
REMEDY
Excessive vibration caused by
loose mounting screws.
Tighten mounting screws.
Excessive tubing vibration.
Tighten clamps and connections,
replace tubing with flexible hose.
TROUBLE SHOOTING -- ALTIMETER.
NOTE
Refer to paragraph 16-15 before blowing out pitot or
static lines.
TROUBLE
INSTRUMENT FAILS TO
OPERATE.
INCORRECT INDICATION.
HAND OSCILLATES.
PROBABLE CAUSE
Static line plugged.
Blow out lines.
Defective mechanism.
Replace instrument.
Hands not carefully set.
Reset hands with knob.
Leaking diaphragm.
Replace instrument.
Pointers out of calibration.
Replace instrument.
Static pressure irregular.
Blow out lines, tighten connections.
Leak in airspeed or vertical
Blow out lines, tighten connections.
speed indicator installations.
16-21.
REMEDY
TROUBLE SHOOTING--VERTICAL SPEED INDICATOR.
NOTE
Refer to paragraph 16-15 before blowing out pitot or
static lines.
TROUBLE
INSTRUMENT FAILS TO
OPERATE.
INCORRECT INDICATION.
PROBABLE CAUSE
REMEDY
Static line plugged.
Blow out lines.
Static line broken.
Repair or replace damaged
line, tighten connections.
Partially plugged static line.
Blow out lines.
Ruptured diaphragm.
Replace instrument.
Pointer off zero.
Reset pointer to zero.
16-9
MODEL 210 & T210 SERIES SERVICE MANUAL
16-21.
TROUBLE SHOOTING (Cont).
POINTER OSCILLATES.
16-22.
REMEDY
PROBABLE CAUSE
TROUBLE
Partially plugged static line.
Blow out lines.
Leak in static line.
Repair or replace damaged lines,
tighten connections.
Leak in instrument case.
Replace instrument.
TROUBLE SHOOTING--PITOT TUBE HEATER.
NOTE
Refer to paragraph 16-15 before blowing out pitot or
static lines.
TUBE DOES NOT HEAT OR
CLEAR ICE.
16-23.
VACUUM SYSTEM.
Switch turned "OFF."
Turn switch "ON."
Popped circuit breaker.
Reset breaker.
Break in wiring.
Repair wiring.
Heating element burned out.
Replace element.
(See figure 16-4.)
16-24. DESCRIPTION. A dry vacuum system is
installed on the aircraft. The system utilizes a
sealed bearing engine-driven vacuum pump. A
discharge tube is connected to the pump to expell
air from the pump overboard. A suction relief valve
is used to control system vacuum and is connected
between the pump inlet and the instruments. A
central air filtering system is utilized. The reading
of the suction gage indicates net difference in suction
before and after air passes through a gyro. This
16-25.
REMEDY
PROBABLE CAUSE
TROUBLE
differential pressure will gradually decrease as the
central air filter becomes dirty, causing a lower
reading on the suction gage. Effective 21064126
barb type fittings are used in the vacuum system to
eliminate the use of hose clamps.
BEGINNING WITH 21064536 a dual pump system is
available. The system plumbing and installation is
illustrated in figure 16-4 sheets 2 of 3 and 3 of 3.
With this system dual vacuum relief valves are
utilized. Both are mounted at Station 3. 85, and
right or left buttock lines 8. 35.
TROUBLE SHOOTING -- VACUUM SYSTEM.
TROUBLE
REMEDY
PROBABLE CAUSE
HIGH SUCTION GAGE READINGS.
(Gyros function normally.)
Relief valve filter clogged,
relief valve malfunction.
Replace filter, reset valve.
Replace gage.
LOW SUCTION GAGE READINGS.
Leaks or restriction between
instruments and relief valve,
relief valve out of adjustment,
defective pump.
Repair or replace lines, adjust or
replace relief valve, repair or replace pump.
Central air filter dirty.
16-10
Revision 1
-
Clean or replace filter.
MODEL 210 & T210 SERIES SERVICE MANUAL
16-25. TROUBLE SHOOTING (Cont).
TROUBLE
SUCTION GAGE FLUCTUATES.
PROBABLE CAUSE
Defective gage or sticking
relief valve.
REMEDY
Replace gage. Clean sticking valve
with Stoddard solvent. Blow dry
and test. If valve sticks after
cleaning, replace it.
16-26.
TROUBLE SHOOTING -- GYROS.
TROUBLE
HORIZON BAR FAILS TO RESPOND.
HORIZON BAR DOES NOT
SETTLE.
HORIZON BAR OSCILLATES OR
VIBRATES EXCESSIVELY.
EXCESSIVE DRIFT IN EITHER
DIRECTION.
PROBABLE CAUSE
REMEDY
Central air filter dirty.
Clean or replace filter.
Suction relief valve improperly adjusted.
Adjust or replace relief valve.
Faulty suction gage.
Replace suction gage.
Vacuum pump failure.
Replace pump.
Vacuum line kinked or
leaking.
Repair or replace damaged lines,
tighten connections.
Defective mechanism.
Replace instrument.
Insufficient vacuum.
Adjust or replace relief valve.
Excessive vibration.
Replace defective shock panel
mounts.
Central air filter dirty.
Clean or replace filter.
Suction relief valve improperly adjusted.
Adjust or replace relief valve.
Faulty suction gage.
Replace suction gage.
Defective mechanism.
Replace instrument.
Excessive vibration.
Replace defective shock panel
mounts.
Central air filter dirty.
Clean or replace filter.
Low vacuum, relief valve improperly adjusted.
Adjust or replace relief valve.
Faulty suction gage.
Replace suction gage.
Vacuum pump failure.
Replace pump.
Vacuum line kinked or
leaking.
Repair or replace damaged lines,
tighten connections.
16-11
MODEL 210 & T210 SERIES SERVICE MANUAL
16-26. TROUBLE SHOOTING GYRO'S (Cont.
TROUBLE
)
PROBABLE CAUSE
DIAL SPINS IN ONE DIRECTION
CONTINUOUSLY.
REMEDY
Operating limits have been
exceeded.
Replace instrument.
Defective mechanism.
Replace instrument.
16-27. TROUBLE SHOOTING -- VACUUM PUMP
TROUBLE
PROBABLE CAUSE
REMEDY
OIL IN DISCHARGE.
Damaged pump drive seal.
Replace gasket.
HIGH SUCTION.
Suction relief valve
filter clogged.
Replace filter.
LOW SUCTION.
Relief valve leaking.
Replace relief valve.
Vacuum pump failure.
Replace vacuum pump.
16-28.
MAINTENANCE PRACTICES.
NOTE
When replacing a vacuum system component,
ensure all connections are made correctly to
avoid damage to gyro system. When a component is removed, cap off and identify all
open lines, hoses, and fittings to prevent dirt
from entering system, and to ensure proper
reinstallation. Upon component replacement,
Check all hoses carefully to be sure they are
clean and free of debris, oil, solvent, collapsed inner liners, and external damage.
Replace old, hard, cracked, or brittle hoses,
particularly on pump inlet, to avoid possible
pump damage. On vacuum pump, where hose
clearance is tight, making it difficult to reinstall hoses, apply a light film of petrolatum
to the fitting. Install hoses by pushing them
straight on, and do not wiggle hoses from side
to side as this could cause particles to be cut
from inside of hose, allowing particles to
enter system.
CAUTION
Do not use teflon tape, pipe dope, or thread
lubricants of any type on fitting threads, and
avoid over-tightening of connections. All
filters in vacuum system must be changed
when installing a new pump. Failure to do
so will void pump warranty. DO NOT CONNECT A PUMP BACKWARDS since the
manifold check valves provide no pressure
relief, the pump will be destroyed within a
matter of seconds after starting the engine.
16-28A. REMOVAL OF VACUUM PUMP.
a. Remove upper engine cowling in accordance with
procedures in Sections 12 of 12A.
b. Disconnect, cap off and identify hose on inlet
side of vacuum pump.
c. Identify and disconnect hose on outlet side of
vacuum pump.
d. Remove nuts, lockwashers, and flat washers
securing vacuum pump to engine.
e. Remove vacuum pump from mounting studs on
engine.
f. Remove elbow from pump and retain if it is reusable.
NOTE
Discard any twisted fittings or nuts with
rounded corners.
16-28B. MOUNTING PAD INSPECTION.
a. Check condition of the AND 20000 pad seal If
the seal shows any signs of oil leakage, replace the
seal. Replace seal if there is any doubt as to its
serviceability.
16-28C. INSTALLATION OF VACUUM PUMP.
a. Before installing a new vacuum pump purge all
lines in the system to remove carbon particles or
pump components that may have been deposited in
the lines by a previous pump.
b. Consult the applicable Parts Catalog. the pump
vendor's application list, or the PMA label on the
pump box to verify that the pump is the correct model
for the engine and/or system.
NOTE
Before installing vacuum pump on engine.
16-12
Revision 2
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
DUAL VACUUM PUMP SYSTEM
(5) is rotated 180 ° clockwise
for clarity.
1.
2.
3.
4.
5.
Gyro Horizon
Directional Gyro
Suction Gage
Central Filter
Manifold Check Valve
Figure 16-4. Vacuum System (Sheet 2 of 3)
16-14
MODEL 210 & T210 SERIES SERVICE MANUAL
7
5
6
1. Right Hand Vacuum Pump
2.
Relief Valve
3. CheckValve Manifold
4. Relief Valve
5.
Left Hand Vacuum Pump
7.
Hose
4
6. Tube
0 BEGINNING WITH 21064773
Figure 16-4. Vacuum System (Sheet 3 of 3)
ensure that mating surfaces are clean and
free of any old gasket material.
c. Position the vacuum pump in a jaw-protected
vise, with drive coupling downward.
__________--~
~Always
{CAUTIOHN
|
Pump housing should never be placed directly
in a vise, since clamping across center housing will cause an internal failure of carbon
rotor. Protect pump mounting flange with soft
metal or wood. NEVER INSTALL a pump that
has been dropped.
NOTE
Do not use teflon tape, pipe dope, or thread
lubricants of any type, and avoid overtightening of connections.
d. Install elbow in pump; hand-tighten only.
NOTE
Use only a box wrench to tighten fittings to
desired position. Do not make more than
one and one half (1-1/2) turns beyond handtighten position.
NOTE
~~~NOTE ~16-29A.
Before
vacuum
nstalling
pump n engine,
ensure that mating surfaces are clean and
feree of any old gasket material
free of any old gaskt
e. Position new mounting pad gasket on mounting
studs on engine.
f. Position vacuum pump on mounting studs.
g. Secure pump to engine with flat washers, new
lockwashers, and nuts.
ICAUTiONl
replace all lockwashers with new ones
when installing a new vacuum pump. Tighten
all four mounting nuts (4) to 50 to 70 poundinches.
h. Connect hose to inlet side of vacuum pump.
i. Install upper engine cowling in accordance with
procedures in Sections 12 or 12A.
16-29. CLEANING. Remove and discard suction
relief valve filter. Wash relief valve with Stoddard
solvent and dry with low pressure, dry compressed
air. Install new filter. Check hoses for external
damage and collapsed inner liners.
[IAUtIONH
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 panelout.
LOW-VACUUMWARNING LIGHT. (See
figure 16-4, sheet 1 of 3.) A red low-vacuum warning light is installed on the instrument panel. This
light is used in conjunction with the single pump
system only. The light is controlled by a vacuum
switch which is teed into the line between the suction
gage and the directional gyro. The switch contacts
Revision 2
16-15
MODEL 210 & T210 SERIES SERVICE MANUAL
are normally closed. The light may be checked by
turning ON the master switch. With the engine running the light should illuminate when the vacuum
drops below 3*. 5 inches Hg.
16-30. VACUUM RELIEF VALVE ADJUSTMENT.
A suction gage reading of 5. 3 inches Hg is desirable
for gyro instruments. However a range of 4.6 to 5.4
inches Hg is acceptable.
Single pump adjustment. Remove central air filter,
run engine at 2200 RPM, adjust relief valve to 5.3*. 1
inches Hg.
Dual pump adjustment. Remove central air filter,
with engine at 1900 set relief valves at lower end of
green arc (4. 8 inches Hg) with individual pump only
on the line. Combined reading (both pumps on line)
not to exceed 5.4 inches Hg at 1900 RPM.
{CAUTION]
Do not exceed maximum engine temperature.
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 the pointer oscillates, check the 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.
NOTE
Before replacing a tachometer cable in the
housing, coat the lower two thirds with AC
Type ST-640 speedometer cable grease or
Lubriplate No. 11Q. Insert the cable in
housing as far as possible, then slowly rotate cable to make sure it is seated in the
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).
16-34. MANIFOLD PRESSURE/FUEL FLOW INDICATOR.
NOTE
With either a single or dual vacuum pump,
if vacuum drops noticeably after replacing
central air filter, remove and replace
existing filter with a new filter.
~ ENGINE
16-31
INIDICATORS.
16-31. ENGINE
INDICATORS.
16 -32 TACHOMETER.
_16-32. TAC~HOMETER,
ii a m
CRI ta
N. D Te
16-33.
16-33. DESCRIPTION.
The tachometer is a mechanical indicator driven at half crankshaft speed by a flexible shaft. Most tachometer difficulties will be found
in the drive-shaft. To function properly, the shaft
SHOP NOTES:
16-16
Revision 2
16-35. 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 in inches of mercury. The fuel fow indicator is a pressure instrument calibrated in pounds
per hour, indicating appraoxmate pounds 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
vete t atmospheric pressure on standard engine
installat atondto turbocharger outlet pressure on
turbocharged engine installations.
MODEL 210 & T210 SERIES SERVICE MANUAL
16-36.
TROUBLE SHOOTING--MANIFOLD PRESSURE INDICATOR.
TROUBLE
PROBABLE CAUSE
EXCESSIVE ERROR AT EXISTING Pointer shifted.
BAROMETRIC PRESSURE.
Leak in vacuum bellows.
REMEDY
Replace instrument.
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.
Rocket 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.
EXCESSIVE POINTER VIBRATION.
Tight rocker pivot bearings.
Replace instrument.
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.
16-17
MODEL 210 & T210 SERIES SERVICE MANUAL
16-37.
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.
16-38. CYLINDER HEAD TEMPERATURE GAGE.
|
16-39. DESCRIPTION. The temperature sending unit
regulates electrical power through the cylinder head
temperature gage. The gage and sender require little or
no maintenance other than cleaning, making sure lead
is properly supported, and all connections are clean,
tight, and properly insulated. Rochester and Stewart
Warner gages are connected the same but the Rochester
gage does not have a calibration pot and cannot be
adjusted. Refer to Table 2 on page 16-22A when trouble
shooting the cylinder head temperature gage.
NOTE
Torque used to tighten wire lead nut not
to exceed 4 inch-pounds.
16-40. TROUBLE SHOOTING.
TROUBLE
GAGE INOPERATIVE.
PROBABLE CAUSE
REMEDY
No current to circuit.
Repair electrical circuit.
Defective gage or sender.
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 voltage supply.
Gage off calibration.
Replace gage or sender.
Check ground connection.
16-18
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
16-40.
TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
GAGE READS TOO LOW
ON SCALE.
REMEDY
Low voltage.
Check voltage supply and
"D" terminal.
Gage off calibration.
Replace defective items.
Defective gage or sender.
Replace defective items.
Defective gage or sender.
Replace defective items.
AT HIGH END.
OBVIOUSLY INCORRECT
READING. READING. ~Incorrect
GAGE READS FULL SCALE
WITH ENGINE COOL OR COLD.
(21064064 & ON)
GAGE READS ZERO WHEN
ENGINE IS HOT.
(21064064 & ON)
16-41.
calibration.
Wire between sender and gage
grounded,
Repair or replace wire as
required.
Defective gage or sender.
Replace defective items.
Wire between gage and sender
is open or disconnected,
Repair or replace wire as
required.
Defective gage or sender.
Replace defective items.
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
16-43.
Replace defective items.
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.
TROUBLE SHOOTING.
TROUBLE
GAGE DOES NOT REGISTER.
GAGE POINTER FAILS TO
RETURN TO ZERO.
PROBABLE CAUSE
REMEDY
Pressure line clogged.
Clean line.
Pressure line broken.
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.
Clean line.
Foreign matter in Bourdon
tube.
Replace instrument.
Bourdon tube stretched.
Replace instrument.
16-19
MODEL 210 &T210 SERIES SERVICE MANUAL
16-43.
TROUBLE SHOOTING. (Cont).
TROUBLE
REMEDY
PROBABLE CAUSE
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.
Repair or replace damaged
line.
16-44.
OIL TEMPERATURE GAGE.
16-45. DESCRIPTION. On some airplanes, the oil
temperature gage is a Bourdon tube type 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 tube's 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.
Refer to Table 1 on page 16-22A when trouble shooting
the oil temperature gage.
16-46.
FUEL QUANTITY INDICATING SYSTEM.
(THRU 21062273).
16-47. INDICATORS. Two fuel quantity indicators,
graduated in pounds/gallons are located in the instrument
cluster. These electromagnetic type indicators are used in
conjunction with a control monitor and capacitance type
sensing units. Refer to paragraph 16-8 for removal and
installation of indicators.
16-48. SENDING UNITS. Two fuel quantity sending
units are located in each fuel bay. These sending units are
basically tubular capacitors with two electrodes fixed in
one position. Any change in fuel quantity between full and
empty produces a corresponding change in the capacitance
of the electrodes. These changes in capacitance are
amplified by the control monitor and actuates the fuel
quantity indicators.
16-48A. REMOVAL AND INSTALLATION. (Refer to
figure 13-2.)
a. Completely drain all fuel from wing bays at bay
sump drain valves. (Observe precautions in Section 13,
Paragraph 13-3.)
b. Remove plates on top of wing bays for access to
sensing units. (Refer to Section 13.)
c. Remove safety wire from probe clips.
16-20
Revision 3
d. Disconnect probe electrical connections and lift
probe out.
e. Reverse the preceding steps for installation.
Prior to reinstalling access plates, calibrate system in
accordance with procedures outlined in paragraph 16-51.
CAUTION
Access plates must be resealed after removal. Refer to Section 13 for sealing
instructions.
16-49. CONTROL MONITOR. The control monitor is
located above the right cabin door, behind the headliner.
A zipper is installed in the headliner for easy access. The
monitor incorporates adjustment provisions for system
calibration.
16-50. REMOVAL AND INSTALLATION.
a. Open zipper in headliner above right door and
remove insulation as necessary.
b. Disconnect all wiring and tag connections for
reference on installation.
c. Remove mounting screws and remove monitor.
d. Reverse preceding steps for installation and
calibrate system in accordance with paragraph 16-51.
16-51.
CALIBRATION.
NOTE
Use field fuel quantity system test box,
PN 2548H, which is available from Barfield
(phone: 800-321-1039). This test box is sold
with an operating instructions manual, or one
may be purchased separately. The field
calibration test box, formerly Cessna
PN 9910111-10, is no longer available.
16-52.
TROUBLE SHOOTING.
NOTE
For additional trouble shooting and testing,
use field fuel quantity system test box,
PN 2548H, which is available from Barfield
(phone: 800-321-1039) and comes with an
operating instructions manual. The field
calibration test box, formerly Cessna
PN 9910111-10, is no longer available.
MODEL 210 & T210 SERIES SERVICE MANUAL
16-52. TROUBLE SHOOTING (Cont).
TROUBLE
NO FUEL QUANTITY
INDICATION.
16-52A.
PROBABLE CAUSE
Fuel bays empty.
Service with proper grade and
amount of fuel.
Circuit breaker open or
defective.
Reset.
Defective fuel quantity
indicator or sending
unit.
Substitute known-good indicator
or sending unit. Replace the
instrument if defective.
Loose connections or open
circuit.
Tighten connections; repair
or replace wiring.
FUEL QUANTITY INDICATING SYSTEM.
(BEGINNING 21062274)
16-52B. DESCRIPTION. The magnetic type fuel quantity indicators are used in conjunction with a floatoperated variable-resistance transmitter in each
fuel tank. The full position of float produces.a mini|
16-52C.
REMEDY
Replace if defective.
mum 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.
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.
(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 voltage.
Revision 3
16-21
MODEL 210 & T210 SERIES SERVICE MANUAL
16-52C. TROUBLE SHOOTING. (Cont.)
TROUBLE
PROBABLE CAUSE
ERRATIC READINGS.
REMEDY
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.
16-52D. REMOVAL AND INSTALLATION. (Refer to
figure 13-2.)
a. Remove access plates on the underside of wing
forward of the flap bellcrank.
b. Drain enough fuel from bay to lower fuel level
below transmitter. (Observe precautions in paragraph 12-3.)
c. Disconnect electrical lead and ground strap
from transmitter.
d. Remove safety wire from transmitter attaching
bolts, remove bolts and carefully remove transmitter
from fuel spar, DO NOT BEND FLOAT ARM.
e. To install transmitter, reverse preceding steps,
using a new gasket around opening in fuel bay and new
sealing washers.
NOTE
16-52E.
TRANSMITTER CALIBRATION.
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.
16-52F. STEWARTWARNERGAGE
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.
Insure that transmitter is grounded per
figure 16-4A.
f.
WARNING
Service fuel bay. Check for leaks and correct
Use extreme caution while working with
fuel quantity indication.
electrical components of the fuel system.
The possibility of electrical sparks
~1.
~around FuelTransmitter
an "empty" fuel cell creates a
hazardous situation.
2.
2. Safety
Safety Wire
Wire
3. Aft Fuel Spar
4. Ground Strap
4. Ground
Upper Wing
Strap
5.
4
g
Before installing transmitter, attach electrical wires and
place master switch "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
16-52G. ROCHESTER GAGE TRANSMITTER. Do
not attempt to adjust float arm or stop. No adjustment is
allowed.
Figure 16-4A.
16-22
Ground Strap Installation
Revision 3
MODEL 210 &T210 SERIES SERVICE MANUAL
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
165º F
120°F
72ºF
250ºF
220°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
46.4
192.0
34.0
990.0
Table 2
NOTE
Select the cylinder head 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
475F
220F
450F
CHT
310.0
34.8
S1372-2
CHT
310.0
34.8
S1372-3
CHT
113.0
S1372-4
CHT
113.0
S2334-3
CHT
745.0
38.0
S2334-4
CHT
745.0
38.0
Part Number
Type
S1372-1
200°F
Revision 3
16-22A/(16-22B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
15-51C.
FUEL QUANTITY INDICATING SYSTEM OPERATIONAL TEST
A. For airplane serials 21061574 thru 21062273:
WARNING:
REMOVE ALL IGNITION SOURCES FROM THE AIRPLANE AND VAPOR HAZARD
AREA. SOME TYPICAL EXAMPLES OF IGNITION SOURCES ARE STATIC
ELECTRICITY, ELECTRICALLY 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. With the fuel selector valve in the "OFF" position, add unusable fuel to each fuel tank.
5. 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.
6. 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.
7.
Install any items and/or equipment removed to accomplish this procedure, remove maintenance
warning tags and connect the airplane battery.
B. For airplane serials 21062274 thru 21064897:
WARNING:
REMOVE ALL IGNITION SOURCES FROM THE AIRPLANE AND VAPOR HAZARD
AREA. SOME TYPICAL EXAMPLES OF IGNITION SOURCES ARE STATIC
ELECTRICITY, ELECTRICALLY 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.
Temporary Revision Number 7
7 October 2002
© 2002 Cessna Aircraft Company
16-22C
MODEL 210 & T210 SERIES SERVICE MANUAL
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
16-52F for instructions to calibrate a Stewart Warner fuel indicating system.
Rochester fuel quantity indicating system components are not adjustable, only
component replacement or standard electrical wiring system maintenance practices are
permitted.
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
16-52F for instructions to calibrate a Stewart Warner fuel indicating system.
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
16-52F for instructions to calibrate a Stewart Warner fuel indicating system.
Rochester fuel quantity indicating system components are not adjustable, only
component replacement or standard electrical wiring system maintenance practices are
permitted.
8.
Install any items and/or equipment removed to accomplish this procedure, remove maintenance
warning tags and connect the airplane battery.
*~~~~~~~~~~~~~ ~~~Temporary
16-22D
©2002 Cessna Aircraft Company
Revision Number 7
7 October 2002
MODEL 210 & T210 SERIES SERVICE MANUAL
16-53.
16-55. ECONOMY MIXTURE INDICATOR.
HOURMETER.
16-54. DESCRIPTION. The hourmeter is an electtrtcally 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 the master switch. A diode incorporated into the
meter prevents interruption of avionics operation.
This type hourmeter is identified by a white + above
the positive terminal.
16-56. DESCRIPTION. The economy mixture indicator is an exhaust gas temperature (EGT) sensing
device which is used to aid the pilot in selecting the
most desirable fuel-air mixture for cruising flight
at less than 75% power. Exhaust gas temperature
(EGT) varies with ratio of fuel-to-air mixture ente.ing the engine cylinders. Refer to the Pilot's Operating Handbook for operating procedure of the syst. .1.
NOTE
When installing the hourmeter, the positive
(red) wire must be connected to the white *
terminal. Connecting wires incorrectly
will damage the meter.
16-57. TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
REMEDY
GAGE INOPERATIVE
Defective gage, probe or
circuit.
Repair or replace defective
part.
INCORRECT READING.
Indicator needs calibrating.
Calibrate indicator in accordance
with paragraph 16-57.
FLUCTUATING READING.
Loose, frayed or broken
lead, permitting alternate
make and break of circuit.
Tighten connections and repair or replace defective
leads.
16-58. CALIBRATION. A potentiometer adjustment
screw is provided either on the front or 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
75% power in level flight, then carefully lean the mixture to peak EGT. After the pointer has peaked,
using the adjustment screw, position pointer over
reference increment line (4/5 of scale).
SHOP NOTES:
16-23
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
16-59. 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.
16-60. MAGNETIC COMPASS.
16-5.)
(Refer to figure
16-61. 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 instrument lights rheostat switch. 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 wind16-66.
shield 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.
16-62. STALL WARNING HORN AND TRANSMITTER.
16-63. DESCRIPTION. The stall warning horn is
contained in the dual warning unit mounted on the
right hand wing root rib. It is electrically operated
and controlled by a stall warning transmitter mounted on the leading edge of the left wing. For further
information on the warning horn and transmitter,
refer to Section 17.
16-64. TURN COORDINATOR.
16-65. DESCRIPTION. The turn coordinator is an
electrically operated, gyroscopic, roll-turn rate
indicator. Its gyro simultaneously senses rate of
motion roll and yaw axis which is projected on a
single indicator. The gyro is a non-tumbling type
requiring no caging mechanism and incorporates an
ac brushless spin motor with a solid state inverter.
TROUBLE SHOOTING.
TROUBLE
INDICATOR DOES NOT RETURN TO CENTER.
PROBABLE CAUSE
REMEDY
Friction caused by contamination
in the indicator dampening.
Replace instrument.
Friction in gimbal assembly.
Replace instrument.
DOES NOT INDICATE A
STANDARD RATE TURN
(TOO SLOW).
Low voltage.
Correct voltage.
Inverter frequency changed.
Replace instrument.
NOISY MOTOR.
Faulty bearings.
Replace instrument.
ROTOR DOES NOT START.
Faulty electrical connection.
Correct voltage or replace
faulty wire.
Inverter malfunctioning.
Replace instrument.
Motor shorted.
Replace instrument.
Bearings frozen.
Replace instrument.
16-24
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
16-66. TROUBLE SHOOTING (Cont).
TROUBLE
IN COLD TEMPERATURES,
HAND FAILS TO RESPOND
OR IS SLUGGISH.
PROBABLE CAUSE
REMEDY
Oil in indicator becomes
too thick.
Replace instrument.
Insufficient bearing end play.
Replace instrument.
Low voltage.
Correct voltage.
16-67. TURN-AND-SLIP INDICATOR.
16-68. DESCRIPTION. The turn-and-slip indicator
isoperated by the aircraft electrical system and
operates ONLY when the master switch is on. Its
circuit is protected by an automatically-resetting
circuit breaker.
16-69. TROUBLE SHOOTING.
TROUBLE
INDICATOR POINTER
FAILS TO RESPOND.
PROBABLE CAUSE
REMEDY
Automatic resetting circuit
breaker defective.
Replace circuit breaker.
Master switch "OFF" or
switch defective.
Replace defective switch.
Broken or grounded lead to
indicator.
Repair or replace defective
wiring.
Indicator not grounded.
Repair or replace defective wire.
Defective mechanism.
Replace instrument.
Defective mechanism.
Replace instrument.
Low voltage.
Correct voltage.
POINTER DOES NOT INDICATE PROPER TURN.
Defective mechanism.
Replace instrument.
HAND DOES NOT SIT ON
ZERO.
Gimbal and rotor out 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.
Correct voltage.
HAND SLUGGISH IN RETURNING TO ZERO.
IN COLD TEMPERATURES,
HAND FAILS TO RESPOND
OR IS SLUGGISH.
16-26
MODEL 210 & T210 SERIES SERVICE MANUAL
16-69.
TROUBLE SHOOTING (Cont).
TROUBLE
NOISY GYRO.
PROBABLY CAUSE
High voltage.
Correct voltage.
Loose or defective rotor
bearings.
Replace instrument.
16-70. ELECTRIC CLOCK.
16-71. DESCRIPTION. The electric clock is connected to the battery through a one-ampere fuse
mounted adjacent to the battery box. The electrical
circuit is separate from the aircraft electrical syster and will operate when the master switch is "OFF."
Beginning with 21062955 a digital clock may be installed. Refer to Pilots Operating Handbook for operating
instructions.
16-72.
REMEDY
FUEL COMPUTER/DIGITAL CLOCK.
16-73. DESCRIPTION. The Astro Tech FT-2 is a
dual function instrument providing a complete fuel
management system and a multi-purpose time keeping device in a single instrument with each function
sharing a common display panel The instrument
may be used as a replacement for the digital or electric clock, and may be mounted in the same location
on the instrument paneL
The fuel computer portion of the instrument displays
the following selections; fuel flow as measured by
an engine mounted transducer, total fuel used, current fuel remaining and time remaining based on fuel
remaining at the current flow rate. Fuel quantities
are displayed in pounds with a gallon display available by utilizing a push button located below and to
the right of the display. When time remaining at the
currect flow rate reaches 45 minutes or less, the
display will be blanked from one-tenth to threetenths of a second per second in all of the selections.
The digital clock portion of the instrument displays
the following selections; current time of day in either
local (LCL) or Greenwich Mean Time (GMT) in hours
and minutes, cummulative flight time in minutes and
seconds (first hour) and hours and minutes (up to 100
hours) whenever fuel flow is greater than 25 to 30
pounds per hour (PPH) and elapsed time in minutes
and seconds (first hour) and hours and minutes (up to
100 hours).
Fuel selections and time selections are made by utilizing a rotary-type selector switch common to both
functions. Two pushbuttons, located below the display, are used to program the fuel computer digital
clock.
16-74. FUEL COMPUTER OPERATION. The fuel
computer contains five selections. They are selected
by rotating the selector switch to the positions labeled ADD, FLOW, LB USD, LB REM, and TIME REM.
These selections, when used in proper sequence with
the programming buttons, will correctly program the
computer.
The fuel quantity added during servicing of the airplane must be entered in the computer so that the
LB REM position accurately represents the correct
amount of usable fuel on board for each flight. The
fuel quantity added is entered in the computer as
follows:
To enter fill-up:
a. Rotate the selector switch to the ADD position.
b. Press left and right programming buttons together until display panel reads FULL.
c. Rotate the selector switch to LB REM position
to display the usable fuel quantity in pounds on board.
NOTE
The usable fuel quantity for each airplane is
programmed into the instrument at the factory.
A
battery
disconnect
other power interruption
will not
alter thisorquantity.
To enter less than fill-up:
a. Rotate the selector switch to the ADD position.
b. Press right programming button, labeled GAL,
until the right digit represents the correct units of
gallons of fuel added.
c. Press left programming button, labeled RST,
until the left two digits represent the correct tens and
hundreds of gallons of fuel added.
d. Rotate the selector switch to LB REM position
to display the correct usable fuel quantity in pounds
on board.
If an error has been made, resulting in an incorrect
display of LB REM, the correct amount may be entered as follows:
a. Leave the selector switch in the ADD position.
b. Enter the corrected fuel quantity in gallons.
c. Rotate the selector switch to FLOW, then press
and hold the left programming button.
d. While holding the left button pressed, slowly
rotate the selector switch to the LB REM position.
The set-in amount in gallons, multiplied by six, will
now appear as LB REM.
When the selector switch is placed in the FLOW position, the display indicates the current fuel flow rate
in pounds per hour (PPH). Press the GAL programming button to display the fow rate in gallons per
hour (GPH).
16-27
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
diameter tool. The reset switch is in a small
diameter hole located between the words "EL
TIME" and "FLT TIME" near the outer periphery of the instrument face. The instrument
should now operate normally, but will have to
be reprogrammed.
16-75. DIGITAL CLOCK OPERATION. The digital
clock contains four selections. They are selected
by rotating the selector switch to the positions labeled
SET, EL TIME, FLT TIME, and LCL/GMT. These
selections, when used, in proper sequence with the
programming buttons, will correctly program the
digital clock.
NOTE
Some models may have an unmarked detent
position between the ADD and SET positions.
This position performs the same function
as the SET position.
.
The digital clock may be set to the local (LCL) and
Greenwich Mean Time (GMT) as follows:
a. Rotate the selector switch to the SET position.
b. Press the left programming button until local
hours advance to the correct value.
c. Press both programming buttons together until
Greenwich Mean Time hours advance to the correct
value.
d. Press right programming button until minutes
advance to correct value. This action sets and holds
seconds to zero.
e. Rotate selector switch from SET to start
seconds from zero hold.
To display the local time-of-day in hours and minutes,
rotate the selector switch to LCL/GMT. If a minutes
and seconds display is desired, press the right programming button, labeled SEC. If Greenwich Mean
Time in hours and minutes is desired, press the left
programming button, labeled GMT.
NOTE
Local or Greenwich Mean Time hours may be
changed without resetting the minutes and
seconds.
To display accumulated flight time, rotate the selector switch to FLT TIME. After the first hour, if a
minutes and seconds display is desired in place of
the hours and minutes display, press the right (SEC)
programming button. Flight time may be reset to
zero by pressing the left (RST) programming button.
NOTE
Accumulated flight time may be zeroed only
when the instrument is not counting (whenever
fuel flow is less than 25-30 PPH) to prevent
accidently zeroing flight time in the air.
Elapsed time (since pressing the RST button) is displayed by rotating the selector switch to the EL
TIME position. After the first hour, if a minutes
and seconds display is desired in place of the hours
and minutes display, press the right (SEC) programming button. Elapsed time may be reset to zero by
pressing the left (RST) programming button.
16-76. TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
REMEDY
Faulty wiring from transducer
to instrument.
Repair or replace wiring.
Faulty transducer
Replace transducer
NO DISPLAY
Faulty wiring or open fuse.
Repair or replace wiring.
Replace fuse.
DISPLAY WILL NOT CHANGE
WITH SELECTOR SWITCH
SELECTION
Low voltage or power
interruption.
Correct low voltage condition.
Connect power supply.
FUEL COMPUTER FUNCTION
INOPERATIVE
Depress reset switch to reset
instrument.
16-29
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
or tabs UP and the turbine totally immersed in fuel.
NOTE
d. Disconnect the electrical connector, connecting
the transducer to the instrument.
e. Disconnect and cap both fuel lines (1 and 7).
Whenever a transducer is installed it must
Remove nuts (5), washers (4), bolts (9) and
calibration
calibration procedures.
procedures.
g.
Reverse these steps for reinstallation.
16-79. TRANSDUCER REMOVAL AND REPLACEMENT (See figure 16-7).
NOTE
When replacing the inlet and outlet pipe fittings
they are to be turned 3 times past hand tight
or torqued to 25-30 lbs-ft whichever occurs
first.
CAUTION
When performing any maintenance on the
fuel system, the precautions in Section 13
must be observed.
The transducer must be mounted horizontally
with the electrical leads on top.
a. Place the fuel selector in the OFF position.
c. Remove the fuse from the clock fuse holder
mounted on the battery contactor bracket.
TOP VIEW
SIDE VIEW
*Torque to 25-30 Lbs/Ft.
TRANSDUCER
*1
SIDE VIEW
~1
c This letter determines the specific setting of
the 3 switches on the back of the fuel computer/
digital clock.
REAR VIEW
FUEL COMPUTER/DIGITAL CLOCK
1. Fuel Computer/Digital Clock
2. Fuel Computer/Digital Clock Switches
3. Transducer
4. Wire Leads
Figure 16-8.
* As an example, the setting shown on the fuel
computer/digital clock switches (2) would be
correct if the boss on top of the transducer
(3) had an "F" stamped on it.
Transducer Markings and Fuel Computer/Digital Clock Switches.
Revision 2
16-31
MODEL 210 & T210 SERIES SERVICE MANUAL
16-80. FUEL TRANSDUCER CALIBRATION. (See
figures 16-8 and 16-9.)
The fuel computer/digital clock (1) has a 3-section switch (2) located on
the back of the unit under a tape cover. Remove
the cover and set the switches as shown on the fuel
transducer table, figure 16-9. The fuel transducer (3) may have one or two letters (stamped or
raised), located on the boss adjacent to the inlet
16-32
Revision 2
port. if the boss contains two letters, DISREGARD
the first letter. The second letter, near the mounting bolt hole, is the calibration "I" factor letter
and determines the switch setting on the fuel conputer/digital clock. After setting the 3 switches
to the transducer marking designation, replace
the tape cover.
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 17
ELECTRICAL SYSTEMS
WARNINGI
When performing any inspection or maintenance
that requires turning on the master switch,
installing a battery, or pulling the propeller
through by hand, treat the propeller as if the
ignition switch were ON. Do not stand nor allow
anyone else to stand, within the arc of the propeller,
since a loose or broken wire or a component
malfunction could cause the propeller to rotate.
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
. 3A4/17-3
ELECTRICAL SYSTEMS ......
3A4/17-3
.
.
....
.
General .
Electrical Power Supply System . . 3A4/17-3
3A4/17-3
Description .........
3A4/17-3
Split Bus Bar .......
3A4/17-3
Description .......
Removal and Installation. . 3A4/17-3
3A4/17-3
Master Switch ........
3A4/17-3
Description ......
3A4/17-3
Ammeter .........
3A4/17-3
......
Description
. .3A4/17-3
Battery Power System ..
3A4/17-3
.....
Battery ...
3A4/17-3
......
Description
3A5/17-4
.....
Trouble Shooting
Removal and Installation. . 3A10/17-9
3A10/17-9
Cleaning the Battery . .
Adding Electrolyte or
Water to the Battery . . 3A10/17-9
3A10/17-9
Testing the Battery ....
Charging the Battery . . .3A10/17-9
3A11/17-10
Battery Box .........
...... 3A11/17-10
Description
Removal and Installation. . 3A11/17-10
3A11/17-10
Maintenance .......
...... 3A11/17-10
Battery Contactor
.......3A11/17-10
Description
3A11/17-10
Removal and Installation.
Battery Contactor Closing
3A11/17-10
Circuit ..........
.3A11/17-10
Ground Service Receptacle
..... 3A11/17-10
Description
.... 3A20/17-19
Trouble Shooting .
Removal and Installation. .3A21/17-20
. 3A21/17-20
Alternator Power System ..
3A21/17-20
.........
Description
. .3A21/17-20
.......
Alternator ..
......
3A21/17-20
Description .
Alternator Reverse Volt
3A20/17-20
...
.
Damage .
. . . 3B1/17-24
Trouble Shooting .
Removal and Installation. .3B6/17-29
3B7/17-30
Alternator Voltage Regulator
3B7/17-30
Description .........
3B7/17-30
Removal and Installation
Alternator Control Unit
(Beginning with 1979 Models) 3B7/17-30
3B7/17-30
Description .............
3B7/17-30
Removal and Installation
Over-Voltage Sensor and
3B7/17-30
Warning Light .......
3B7/17-30
Description .......
Removal and Installation. . 3B7/17-30
Rigging Throttle-Operated
. 3B7/17-30
Microswitch ......
Auxiliary Fuel Pump Flow Rate
3B7/17-30
.....
Adjustment .
. 3B11/17-34
Standby Generator System.
.......
3B11/17-34
Description
3B11/17-34
Removal and Installation.
. . . 3B11/17-34
Dual Alternator System
3B11/17-34
Description .......
3B11/17-34
Alternators .........
3B11/17-34
Description .......
Removal and Installation. . 3B11/17-34
Alternator Control Units . . . 3B11/17-34
3B11/17-34
Description .......
Removal and Installation. . 3B11/17-34
Alternator Contactors and
3B11/17-34
Shunts ...........
3B11/17-34
Description .......
Removal and Installation. . 3B11/17-34
3B11/17-34 I
Volt-Ammeter ........
3B11/17-34
Description .......
Alternator Restart System. . . 3B11/17-34
3B21/17-44
Aircraft Lighting System .....
3B21/17-44
Description .........
3B21/17-44
Switches ........
3B21/17-44
Description .......
3B21/17-44
Trouble Shooting .......
Landing and Taxi Lights . . . 3C1/17-48
3C1/17-48
Description .......
Removal and Installation. . 3C1/17-48
3C1/17-48
Navigation Lights. .......
3C1/17-48
Description .......
Removal and Installation. .3C1/17-48
Anti-Collison Strobe Lights .. 3C1/17-48
3C1/17-48
Description ............
Operational Requirements
3C1/17-48
(Thru 1977 Models) ....
3C3/17-50
Removaland nstallation
3C3/17-50
Vertical Tail Flood Lights
3C3/17-50
Description.
3C3/17-50
Removal and Installation
3C3/1750
FlashingBeacon ............
3C3/17-50
Description ...........
3C3/17-50
Removal and Installation
3C3/17-50
Instrument Lighting ........
Revision 3
17-1
17-2
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
17-1.
ELECTRICAL SYSTEMS.
17-2. GENERAL. This section contains service information necessary to maintain the Aircraft Electrical Power Supply System, Battery and External
Power Supply System, Alternator Power System, Aircraft Lighting System, Pitot Heater, Stall Warning,
Cigar Lighter, and Electrical Load Analysis.
17-3.
ELECTRICAL POWER SUPPLY SYSTEM.
17-4. DESCRIPTION. Energy for the aircraft is
supplied by a 28- volt, direct-current, single wire,
negative ground electrical system. A 24-volt battery
supplies power for starting and furnishes a reserve
in event of alternator failure. An alternator is the
normal source of power during flight and maintains
a battery charge controlled by a voltage regulator,
An external power source receptacle may be installed to supplement the battery alternator system for
starting and ground operation.
17-5. SPLIT BUS BAR.
17-6. DESCRIPTION. Electrical power is supplied
through two bus bars. Thru 1977 Models one bus bar
is located on the lower left hand side of the instrument
panel. This bus bar supplies power to the electrical
equipment. The other bus bar powers the electronic
equipment, and is located on the left hand cabin side
forward of the cabin door. Beginning with 1978 Models
both bus bars are located on the cabin side forward
of the left hand door. A avionics master switch is installed on the electronic bus bar to prevent transient
voltages from damaging the semiconductor circuitary
in the electronic installations.
17-7. REMOVAL AND INSTALLATION. (Refer to
17-8.
MASTER SWITCH.
17-9. DESCRIPTION. The operation of the battery
and alternator systems is controlled by a master
switch. The switch is an interlocking split rocker
with the battery mode on the right-hand side and the
alternator mode on the left-hand side. This arrangement allows the battery to be on the line without the
alternator, however, operation of the alternator without the battery on the line is not possible. The switch
is labeled "BAT" and "ALT" below the switch and is
located on the left-hand side of the switch panel.
17-10. AMMETER.
17-11. DESCRIPTION. The ammeter is connected
between the battery and the aircraft bus. The meter
indicates the amount of current flowing either to or
from the battery. With a low battery and the engine
operating at cruise speed the ammeter will show the
full alternator output when all electrical equipment is
off. When the battery is fully charged and cruise
RPM is maintained with all electrical equipment off,
the ammeter will show a minimum charging rate.
17-12. BATTERY POWER SYSTEM.
17-13. BATTERY.
17-14. DESCRIPTION. The battery is 24 volts and
thru 21062273 a 14 ampere-hour capacity battery is
installed as standard, a 17 ampere-hour capacity
battery is optional. Beginning with 21062274 the battery is 24 volts with a 12.75 ampere-hour capacity
as standard and a 15.5 ampere-hour capacity battery
as optional. The battery is mounted on the forward
left side of the firewall and is equipped with non-spill
caps.
figure 17-1. )
17-3
MODEL 210 & T210 SERIES SERVICE MANUAL
17-15. TROUBLE SHOOTING.
TROUBLE
BATTERY WILL NOT SUPPLY
POWER TO BUS OR IS INCAPABLE OF CRANKING ENGINE
PROBABLE CAUSE
Battery discharged.
1. Measure voltage at "BAT"
terminal of battery contactor
with master switch and a suitable load such as a taxi light
turned on. Normal battery will
indicate 23 volts. If voltage is
low proceed to step 2. If voltage is normal proceed to step
3.
Battery faulty.
2. Check fluid level in cells
and charge at 28 volts for approximately 30 minutes or until battery voltage rises to 28
volts. If tester indicates a good
battery, the malfunction may
be assumed to be a discharged
battery. If tester indicates a
faulty battery, replace the
battery.
Faulty contactor or wiring.
between contactor and master
switch.
3. Measure voltage at master
switch terminal (smallest) on
contactor with master switch
closed. Normal indication is
zero volts. If voltage reads
zero, proceed to step 4. If a
voltage reading is obtained,
check wiring between contactor
and master switch. Also check
master switch.
4. Check continuity between
"BAT" terminal and master
switch terminal of-contactor.
Normal indication is 50-70
ohms. If ohmmeter indicates
an open coil, replace contactor.
If ohmmeter indicates a good
coil, proceed to step 5.
Open coil on contactor.
17-4
REMEDY
Faulty contactor contacts.
5. Check voltage on "BUS"
side of contactor with master
switch closed. Meter normally indicates battery voltage.
If voltage is zero or intermittent, replace contactor. If
voltage is normal, proceed to
step 6.
Faulty wiring between contactor and bus.
6. Inspect wiring between contactor and bus. Repair or
replace wiring.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
17-16. REMOVAL AND INSTALLATION OF THE
BATTERY. (Refer to figure 17-2).
a. To gain access to the battery, remove the upper
left half of cowling.
b. Remove the battery box lid and disconnect the
battery ground cable.
CAUTION
Always remove the ground cable first and connect it last to prevent accidentally shorting the
battery to the airframe with tools.
c. Disconnect the positive cable from the battery
and remove the battery from the aircraft.
d. To install a battery, reverse this procedure.
17-17. CLEANING THE BATTERY. For maximum
efficiency, the battery and connections should be kept
clean at all times.
a. Remove the battery in accordance with preceding paragraph.
b. Tighten battery cell filler caps to prevent the
cleaning solution from entering the cells.
c. Wipe battery cable ends, battery terminals and
entire surface of the battery with a clean cloth moistened with a solution of bicarbonate of soda (baking
soda) and water.
d. Rinse with clear water, wipe off excess water
and allow battery to dry.
e. Brighten up cable ends and battery terminals
with emery cloth or a wire brush.
f. Install the battery according to the preceding
paragraph.
g. Coat the battery terminals and the cable ends
with petroleum jelly.
17-18. ADDING ELECTROLYTE OR WATER TO
THE BATTERY. A battery being charged and discharged with use will decompose the water from the
electrolyte by electrolysis. When the water is decomposed, hydrogen and oxygen gases are formed
which escape into the atmosphere through the battery
vent system. The acid in the solution chemically
combines with the plates of the battery during discharge or is suspended in the electrolyte solution
during charge. Unless the electrolyte has been
spilled from a battery, acid should not be added to
the solution. The water will decompose into gases
and should be replaced regularly. Add distilled
water as necessary to maintain the electrolyte level,
(thru 21062273 the 12-GCAB-9 battery) 3/8 inch
above separators, (beginning with 21062274 the
G-240 and G-242 batteries) to the bottom of split
ring. When "dry charged" batteries are put into
service, fill as directed with electrolyte. However
as the electrolyte level falls below normal with use
add only distilled water to maintain the proper level.
The battery electrolyte contains approximately 25%
sulphuric acid by volume. Any change in this
volume will hamper the proper operation of the
battery.
CAUTION
Do not add any type of "battery rejuvenator"
to the electrolyte. When acid has been
spilled from a battery, the acid balance may be
adjusted by following instructions published
by the Association of American Battery
Manufacturers.
17-19. TESTING THE BATTERY. The specific
gravity check method of testing the battery is preferred when the condition of the battery is in a
questionable state-of-charge. However, when the
aircraft has been operated for a period of time with
an alternator output voltage which is known to be
correct, the question of battery capability may be
answered more correctly with a load type tester. If
testing the battery is deemed necessary, the specific
gravity should be checked first and compared with
the following chart.
BATTERY HYDROMETER READINGS
1.280
1.250
1.220
1. 190
1. 160
Specific
Specific
Specific
Specific
Specific
Gravity
Gravity
Gravity
Gravity
Gravity
100% Charged
75% Charged
50% Charged
25% Charged
Practically Dead
NOTE
All readings shown are for an electrolyte
temperature of 80°F (27°C). For higher
temperatures the readings will be slightly
lower. For cooler temperatures the readings
will be slightly higher. Some hydrometers
have a built-in temperature compensation
chart and a thermometer. If this type tester
is used, disregard this chart.
If the specific gravity reading indicates the battery
is not fully charged the battery should be charged.
The charging rate for the 12-GCAB-9 battery is
2 amps to start and finish at 1 amp, on the G-240
battery, 2 amps and on the G-242 battery, 3 amps.
17-20. CHARGING THE BATTERY. When the battery is to be charged, the level of electrolyte should
be checked and adjusted by adding distilled water to
cover the tops of the internal battery plates. The
battery cables and connections should be clean.
Remove the battery from the aircraft and place
in a well ventilated area for charging.
WARNING
When a battery is charging, hydrogen and
oxygen gases are generated. Accumulation
of these gases can create a hazardous explosive condition. Always keep sparks and
open flame away from the battery. Allow
unrestricted ventilation of the battery area
during charging.
The main points of consideration during a battery
charge are excessive battery temperature and violent
gassing. Under a reasonable rate of charge, the battery
temperature should not rise over 115°F (46°C) (see
paragraph 17-19), nor should gassing be so violent that
acid is blown from the vents.
Revision 3
17-9
MODEL 210 & T210 SERIES SERVICE MANUAL
17-21.
BATTERY BOX.
17-22. DESCRIPTION. The battery is completely
enclosed in a box which is painted with acid proof
paint. The box has a vent tube which protrudes
through the bottom of the aircraft allowing battery
gases and spilled electrolyte to escape. The battery
box is riveted to the left forward side of the firewall.
17-23. REMOVAL AND INSTALLATION. (Refer to
figure 17-2.) The battery box is riveted to the firewall.
The rivets must be drilled out to remove the box. When
a battery box is installed and riveted into place, all
rivets and scratches inside the box should be painted
with acid-proof lacquer, available from Pratt and
Lambert United - Performance Coatings Division, P. O.
Box 2153, Wichita, KS 67201.
17-24. MAINTENANCE. The battery box should be
inspected and cleaned periodically. The box and
cover should be cleaned with a strong solution of
bicarbonate of soda (baking soda) and water. Hard
deposits may be removed with a wire brush. When
all corrosive deposits have been removed from the
box, flush it thoroughly with clean water.
WARNING
Do not allow acid deposits to come in contact
with skin or clothing. Serious acid burns
may result unless the affected area is washed
immediately with soap and water. Clothing
will be ruined upon contact with battery acid.
Inspect the cleaned box and cover for physical damage
and for areas lacking proper acid proofing. A badly
damaged or corroded box should be replaced. If the box
or lid require acid proofing, paint the area with acidproof black lacquer, available from Pratt and Lambert
United - Performance Coatings Division, P. O. Box 2153,
Wichita, KS 67201.
17-25. BATTERY CONTACTOR.
17-26. DESCRIPTION. The battery contactor is
bolted to the firewall below the battery box. The contactor is a solenoid plunger type, which is actuated
by turning the master switch on. When the master
switch is off, the battery is disconnected from the
electrical system. A silicon diode is used to eliminate spiking of the transistorized radio equipment
The cathode (+) terminal of the diode connects to the
battery terminal of the battery contactor. The anode
(-) terminal of the diode connects to the same terminal of the diode connects to the same terminal on the
contactor as the master switch wire. This places the
diode directly across the contactor solenoid coil so
that inductive spikes originating in the coil are clipped
when the master switch is opened. (Refer to figure
17-2).
17-28. BATTERY CONTACTOR CLOSING CIRCUIT.
(Refer to figure 17-3). This circuit consists of a 5amp fuse, a resistor and a diode mounted on the
ground service receptacle bracket. This serves to
shunt a small charge around the battery contactor so
that ground power may be used to close the contactor
when the battery is too low to energize the contactor
by itself.
17-29. GROUND SERVICE RECEPTACLE.
17-30. DESCRIPTION. A ground service receptacle
is installed to permit the use of external power for
cold weather starting or when performing lengthy
electrical maintenance. A reverse polarity protection system is utilized whereby ground power must
pass through an external power contactor to be connected to the bus. A silicon junction diode is connected in series with the coil on the external power contactor so that if the ground power source is inadvertently connected with a reversed polarity, the external power contactor will not close. This feature
protects the diodes in the alternator, and other semiconductor devices used in the aircraft, from possible
reverse polarity damage.
NOTE
Maintenance of the electronic installations
cannot be performed when using external
power. Application of external power opens
the relay supplying voltage to the electronics
bus. For lengthy ground testing of electronic
systems, connect a well regulated and filtered
power supply directly to the battery side of the
battery contactor. Adjust the supply for 28
volts and close the master switch.
NOTE
When using ground power to start aircraft, close
the master witch before removing ground power
plug. This ill ensure closure of battery contactor and excitaton of the alterator field.
CAUTION
Failure to observe polarity when connecting
an external power source directly to the bat-
tery or directly to the battery side of the bat-
17-27. REMOVAL AND INSTALLATION. (Referto
figure 17-2.)
a. Open battery box (2) and disconnect ground cable
(8) from negative battery terminal. Pull cable clear of
battery box.
17-10
b. Remove the nut, lockwasher, and two plain
washers securing the battery cables to the battery
contactor (4).
c. Remove nut, lockwasher, and two plain washers
securing the wire which is routed to the master switch.
d. Remove bolt, washer, and nut securing each side of
the battery contactor (4). The contactor will now be free
for removal.
e. To replace the contactor, reverse this procedures.
Revision 3
tery contactor, will damage the diodes in the
alternator and other semiconductor devices
in the aircraft.
NOTE
On Aircraft Serials 21061574 thru 21062334
On Aircraft Serials 21061574 thru 21062334
refer to Cessna Single-engine Service Letter
SE78-19, dated March 27, 1978.
|
MODEL 210 & T210 SERIES SERVICE MANUAL
9
10
Detail A
(Cover Removed)
THRU 21064135
* BEGINNING WITH 21062274
* BEGINNING WITH 1979 MODELS
Figure 17-2. Battery and Electrical Equipment Installation (Sheet 3 of 5)
17-13
MODEL 210 & T210 SERIES SERVICE MANUAL
29
8 7
14
18
Detail A
28. Jumper Wire
29.
30.
31.
32.
33.
34.
Cover (Battery Contactor)
Cover (Terminal Block)
Terminal Block
Firewall
P C Board
Cover (Starter Contactor)
BEGINNING WITH 21064136
Figure 17-2. Battery and Electrical Equipment Installation (Sheet 4 of 5)
17-14
MODEL 210 & T210 SERIES SERVICE MANUAL
31
MODEL 210 & T210 SERIES SERVICE MANUAL
17-31. TROUBLE SHOOTING.
TROUBLE
GROUND POWER WILL NOT
CRANK ENGINE.
PROBABLE CAUSE
Ground service connector
wired incorrectly.
REMEDY
1. Check for voltage at all
three terminals of external
power contactor with ground
power connected and master
switch off. If voltage is present on input and coil terminals but not on the output terminal. proceed to step 4. If
voltage is present on the input
terminal but not on the coil
terminal, proceed to step 2.
If voltage is present on all three
terminals, check wiring between
contactor and bus.
2. Check for voltage at small
terminal of ground service receptacle. If voltage is not present, check ground service plug
wiring. If voltage is present,
proceed to step 3.
Open or mis-wired diode on
ground service diode board
assembly.
3. Check polarity and continuity
of diode on diode board at rear
of ground service receptacle.- If
diode is open or improperly wired,
replace diode board assembly.
Faulty external power contactor.
4. Check resistance from small
(coil) terminal of external power
contactor to ground (master switch
off and ground power unplugged).
Normal indication is 50-70 ohms
If resistance indicates an open
coil, replace contactor. If resistance is normal, proceed to
step 5.
Faulty contacts in external
power contactor.
5. With master switch off and
ground power applied, check for
voltage drop between two large
terminals of external power
(turn on taxi light for a load).
Normal indication is zero volts.
If voltage is intermittently present or present all the time,
replace contactor.
17-19
MODEL 210 & T210 SERIES SERVICE MANUAL
17-32. REMOVAL AND INSTALLATION. (Refer to
figure 17-3.)
a. Openthe battery box and disconnect the ground
cable from the negative terminal of the battery and
pull the cable free of the box.
b. Remove the nuts, washers, ground strap, bus
bar and diode board from the studs of the receptacle
and remove battery cable
c. Remove the screws and nuts holding the receptacle. ground strap will then be free from bracket.
d. To install a ground service receptacle, reverse
this procedure.
17-33.
ALTERNATOR POWER SYSTEM.
17-34. DESCRIPTION. The alternator system consists of an engine driven alternator, a voltage regulator and a circuit breaker located on the instrument
panel. The system is controlled by the left hand portion of the split rocker, master switch labeled ALT.
An over-voltage sensor switch and red warning light,
labeled HIGH VOLTAGE are incorporated to protect
SHOP NOTES:
17-20
the system. The aircraft battery supplies the source
of power for excitation of the alternator.
17-35. ALTERNATOR.
17-36. DESCRIPTION. The 60-ampere alternator
used on the aircraft is three-phase, delta connected
with integral silicon diode rectifiers. The alternator
is rated at 28-volts at 60-amperes continuous output. Beginning with 1978 Models a 28-volt, 95 ampere
alternator may be installed.
17-37. ALTERNATOR REVERSE VOLTAGE DAMAGE. The alternator is very-susceptible to reverse
polarity damage due to the very low resistance of the
output windings and the low resistance of the silicon
diodes in the output. If a high current source, such
as a battery or heavy duty ground power cart is attached to the aircraft with the polarity inadvertently
reversed, the current through the alternator will
flow almost without limit and the alternator will be
immediately damaged.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
3
* TORQUE TO 165 ± 10 IN LBS.
* TORQUE TO 450 - 500 IN LBS.
BEGINNNG WITH 21062650,
21062662 AND 21062667 & ON
Figure 17-4. Alternator Installation (Sheet 2 of 3)
17-22
MODEL 210 & T210 SERIES SERVICE MANUAL
17-38. TROUBLE SHOOTING THE ALTERNATOR SYSTEM (THRU 1978 MODELS).
a. ENGINE NOT RUNNING.
TROUBLE
PROBABLE CAUSE
REMEDY
AMMETER INDICATES HEAVY
DISCHARGE OR ALTERNATOR
CIRCUIT BREAKER OPENS.
(Battery Switch ON,
Alternator Switch OFF,
all other electrical
switches OFF. )
Shorted diode in alternator.
Turn off Battery Switch and remove "B" Lead from alternator.
Check resistance from "B"
Terminal of alternator to alternator case. Reverse leads and
check again. Resistance reading
may show continuity in one direction but should show an infinite
reading in the other direction.
If an infinite reading is not obtained in at least one direction,
repair or replace alternator.
ALTERNATOR REGULATOR
CIRCUIT BREAKER OPENS
WHEN BATTERY AND
ALTERNATOR SWITCHES
ARE TURNED ON.
Short in Over-Voltage
sensor.
Disconnect Over-Voltage Sensor
plug and recheck. If circuit
breaker stays in replace OverVoltage Sensor.
Short in alternator voltage
regulator.
Disconnect regulator plug and
recheck. If circuit breaker
stays in, replace regulator.
Short in alternator field.
Disconnect "F" terminal wire
and recheck. If circuit breaker
stays in, replace alternator.
ALTERNATOR CIRCUIT
BREAKER OPENS WHEN
BATTERY AND ALTERNA TOR SWITCHES ARE
TURNED ON, OVERVOLTAGE LIGHT DOES
NOT COME ON.
Defective circuit breaker.
Replace circuit breaker.
ALTERNATOR REGULATOR
CIRCUIT BREAKER OPENS
WHEN BATTERY AND
ALTERNATOR SWITCHES
ARE TURNED ON, OVERVOLTAGE LIGHT DOES
NOT COME ON
Shorted field in alternator.
Check resistance from "F"
terminal of alternator to
alternator case, if resistance
is less than 5 ohms repair/
replace.
b.
ENGINE RUNNING.
CAUTION
This malfunction frequently causes a shorted regulator which
will result in an over-voltage condition when system is again
operated.
17-24
MODEL 210 & T210 SERIES SERVICE MANUAL
17-38. TROUBLE SHOOTING THE ALTERNATOR SYSTEM (THRU 1978 MODELS) (Cont.)
b. ENGINE RUNNING (Cont.)
REMEDY
PROBABLE CAUSE
TROUBLE
ALTERNATOR MAKES
ABNORMAL WHINING
NOISE.
Shorted diode in alternator.
Turn off Battery Switch and
remove "13" Lead from
alternator. Check resistance
from "B" Terminal of alternator to alternator case. Reverse leads and check again.
Resistance reading may show
continuity in one direction but
should show an infinite reading
in the other direction. If an
infinite reading is not obtained
in at least one direction, repair
or replace alternator.
OVER-VOLTAGE LIGHT DOES
NOT GO OUT WHEN ALTERNATOR AND BATTERY
SWITCHES ARE TURNED ON.
Shorted regulator.
Replace regulator.
Defective over-voltage
sensor.
Replace sensor.
AFTER ENGINE START
Regulator faulty or high
With engine not running turn
WITH ALL ELECTRICAL
EQUIPMENT TURNED OFF
CHARGE RATE DOES NOT
TAPER OFF IN 1-3 MINUTES
resistance in field circuit.
off all electrical loads and
turn on battery and alternator
switches. Measure bus voltage
to ground, then measure voltage
from terminal of alternator to
ground. If there is more than
2 volts difference check field
circuit wiring shown on alternator system wiring diagram
in Section 19. Clean all contacts.
Replace components until there
is less than 2 volts difference
between bus voltage and field
voltage.
NOTE
Also refer to battery power system trouble shooting chart.
ALTERNATOR SYSTEM WILL
NOT KEEP BATTERY
CHARGED.
Alternator output voltage
insufficient.
1.
Connect voltmeter between
D. C. Bus and ground. Turn off
all electrical loads. Turn on
Battery Switch, start engine and
adjust for 1500 RPM. Voltage
should read approximately 24 volts
Turn on alternator switch, voltage should read between 27.4
and 28.0 volts. Ammeter should
indicate a heavy charge rate which
should taper off in 1-3 minutes.
If charge rate tapers off very
17-25
MODEL 210 & T210 SERIES SERVICE MANUAL
17-38. TROUBLE SHOOTING THE ALTERNATOR SYSTEM (THRU 1978 MODELS) (Cont.)
b. ENGINE RUNNING (Cont.)
TROUBLE
ALTERNATOR SYSTEM WILL
NOT KEEP BATTERY
CHARGED. (Cont.)
PROBABLE CAUSE
Alternator output voltage
insufficient (cont).
REMEDY
quickly and voltage is normal,
check battery for malfunction.
If ammeter shows a low charge
rate or any discharge rate,
and voltage does not rise when
alternator switch is turned on
proceed to Step 2.
2. Stop engine, turn off all
switches. Connect voltmeter
between "F" terminal of
alternator and ground. Do
NOT start engine. Turn on
battery switch and alternator
switch. Battery voltage
should be present at "F"
terminal, less 1 volt drop
thru regulator, if not refer
to Step 3.
3. Starting at "T" terminal
of alternator trace circuit
to voltage regulator, at "B"
terminal of regulator trace
circuit to over-voltage sensor,
to master switch, to Bus Bar.
Replace component which does
not have voltage present at
output. Refer to alternator
system wiring diagram in
Section 19.
Alternator field winding
open.
1. If voltage is present turn off
alternator and battery switches.
Check resistance from "F"
terminal of alternator to alternator case, turning alternator
shaft during measurement.
Normal indication is 12-20 ohms.
If resistance is high or low,
repair or replace alternator. If
ok refer to Step 2.
2. Check resistance from case
of alternator to airframe ground.
Normal indication is very low
resistance. If reading indicates
no, or poor continuity, repair or
replace alternator ground wiring.
17-26
MODEL 210 & T210 SERIES SERVICE MANUAL
17-38A. TROUBLE SHOOTING THE ALTERNATOR SYSTEM (BEGINNING WITH 1979 MODELS).
a. ENGINE NOT RUNNING.
PROBABLE CAUSE
TROUBLE
REMEDY
AMMETER INDICATES
HEAVY DISCHARGE OR
ALTERNATOR CIRCUIT
BREAKER OPENS.
(Battery Switch ON. Alternator Switch OFF. all
other electrical switches
OFF.)
Shorted diode in alternator.
Turn off Battery Switch and
remove "B" Lead from alternator. Check resistance from
"B" Terminal of alternator to
alternator case. Reverse
leads and check again. Resistance reading may show continuity in one direction but
should stow an infinite reading
in the other direction. If an
infinite reading is not obtained
in at least one direction. repair
or replace alternator.
ALTERNATOR REGULATOR CIRCUIT BREAKER
OPENS WHEN BATTERY
AND ALTERNATOR
SWITCHES ARE TURNED
ON.
Short in alternator control
unit.
Disconnect Over-Voltage
Sensor plug and recheck.
If circuit breaker stays in
replace Over-Voltage Sensor.
Disconnect alternator control
unit plug and recheck. If
circuit breaker stays in. replace
alternator control unit.
Short in alternator field.
Disconnect "F" terminal wire
and recheck. If circuit
breaker stays in. replace
alternator
ALTERNATOR CIRCUIT
BREAKER OPENS WHEN
BATTERY AND ALTERNATOR SWITCHES ARE
TURNED ON. LOWVOLTAGE LIGHT DOES
NOT COME ON.
Defective circuit breaker
Replace circuit breaker.
ALTERNATOR REGULATOR CIRCUIT BREAKER
OPENS WHEN BATTERY
AND ALTERNATOR
SWITCHES ARE TURNED
ON, LOW-VOLTAGE
LIGHT MAY OR MAY NOT
COME ON.
Shorted field in alternator.
Check resistance from "F"
terminal of alternator to
alternator case, if resistance is less than 5 ohms
repair/replace.
b.
ENGINE RUNNING.
CAUTION
This malfunction may cause a snorted alternator control unit. which
will result in an over-voltage condition when system is again operated.
17-27
MODEL 210 & T210 SERIES SERVICE MANUAL
17-38A. TROUBLE SHOOTING THE ALTERNATOR SYSTEM (BEGINNING WITH 1979 MODELS) (Cont.)
b. ENGINE RUNNING (Cont.)
TROUBLE
PROBABLE CAUSE
REMEDY
ALTERNATOR MAKES
ABNORMAL WHINING
NOISE.
Shorted diode in alternator.
Turn off Battery Switch and
remove "B" Lead from alternator. Check resistance
from "B" Terminal of alternator to alternator case. Reverse leads and check again.
Resistance reading may show
continuity in one direction but
should show an infinite reading
in the other direction. If an
infinite reading is not obtained
in one direction, repair or
replace alternator.
LOW-VOLTAGE LIGHT
DOES NOT GO OUT WHEN
ALTERNATOR AND BATTERY SWITCHES ARE
TURNED ON.
Shorted alternator control
unit.
Replace alternator control unit.
Defective low-voltage
sensor.
Replace alternator control unit.
AFTER ENGINE START
WITH ALL ELECTRICAL
EQUIPMENT TURNED OFF
CHARGE RATE DOES NOT
TAPER OFF IN 1-3
MINUTES
Alternator control unit faulty
or high resistance in field
circuit
With engine not running turn
off all electrical loads and
turn on battery and alternator
switches. Measure bus voltage to ground. then measure
voltage from terminal of
alternator to ground. If there
is more than 2 volts difference
check field circuit wiring shown
in alternator system wiring
diagram in Section 19 Clean
all contacts. Replace components
until there is less than 2 volts
difference between bus voltage and
field voltage.
NOTE
Also refer to battery power system trouble shooting chart.
ALTERNATOR SYSTEM
WILL NOT KEEP BATTERY CHARGED.
17-28
Alternator output voltage
insufficient.
1. Connect voltmeter between
D. C. Bus and ground. Turn
off all electrical loads. Turn
on Battery Switch, start
engine and adjust for 1500 RPM.
voltage should read approximately
24 volts. Turn on-alternator switch.
voltage should read between 28.4
and 28.9 volts. Ammeter should
indicate a heavy charge rate which
should taper off in 1-3 minutes.
If charge rate tapers off very
quickly and voltage is normal.
check battery for malfunction. If
ammeter shows a low charge rate
or any discharge rate, and voltage
does not rise when alternator
switch is turned on proceed to
Step 2.
MODEL 210 & T210 SERIES SERVICE MANUAL
17-38A. TROUBLE SHOOTING THE ALTERNATOR SYSTEM (BEGINNING WITH 1979 MODELS) (Cont.)
b. ENGINE RUNNING (Cont.)
TROUBLE
ALTERNATOR SYSTEM
WILL NOT KEEP BATTERY CHARGED. (Cont.
PROBABLE CAUSE
Alternator output voltage
insufficient (cont.)
REMEDY
2. Stop engine. turn off all
switches. Connect voltmeter
between "F" terminal of
alternator and ground. Do
NOT start engine. Turn on
battery switch and alternator
switch. Battery voltage
should be present at "F "
terminal, less 1 volt drop
thru regulator, if not refer
to Step A3.
3. Starting at "F" terminal of
alternator, trace circuit to
alternator control unit at
Pin 1 (Blue Wire). Trace
circuit from Pin 3 (Red Wire)
to master switch, to Bus Bar.
Trace circuit from alternator
control unit Pin 2 (Orange Wire)
to alternator "BAT" terminal.
Check connections and replace
component which does not have
voltage present at output. Refer
to alternator system wiring
diagram in Section 19.
Alternator field winding
open.
1. If voltage is present turn
off alternator and battery
switches. Check resistance
from 'F" terminal of alternator to alternator case.
turning alternator shaft during measurement. Normal
indication is 12-20 ohms.
If resistance is high or low.
repair or replace alternator.
If OK refer to Step 2.
Alternator output voltage
insufficient.
2. Check resistance from case
of alternator to airframe ground.
Normal indication is very low
resistance. If reading indicates
no, or poor continuity, repair or
replace alternator ground wiring.
17-39. REMOVAL AND INSTALLATION. (Refer to
figure 17-4, Sheet 3, typical.)
a. Make sure that master switch remains in the off
position, or disconnect negative lead from battery.
b. Disconnect wiring from the alternator.
c. Remove safety wire (4) from the upper adjusting
bolt (3), and remove bolt from alternator.
d. Remove nut (7) and washer (2) from the lower
mounting bolt.
e. Remove alternator drive belt (5) and lower bolt (3)
to remove alternator.
f. To replace alternator, reverse this procedure.
lb-in
g. Adjust belt tension to obtain 3/8-inch deflection at
the center of the belt when applying 12 pounds of
pressure to the belt. After the belt is adjusted and the
bolt is safety wired, tighten the bottom bolt to 100-140
torque on the 60 ampere alternator and 450-500
lb-in torque on the 95 ampere alternator to remove any
play between the alternator mounting foot and the
U-shaped support assembly.
CAUTION
On new aircraft or whenever a now belt is
installed, belt tension should be checked
within 10 to 25 hours of operation.
Revision 3
17-29
MODEL 210 & T210 SERIES SERVICE MANUAL
NOTE
When tightening the alternator belt, apply pry
bar pressure only to the end of the alternator
nearest to the belt pulley.
17-40.
ALTERNATOR VOLTAGE REGULATOR.
1741. DESCRIPTION. A transistorized voltage
regulator is installed on the aircraft. The regulator is
adjustable, but adjustment on the aircraft is not
recommended. A bench adjustment procedure is
outlined in the Cessna Alternator Charging Systems
Service/Parts Manual. A Cessna Alternator Charging
System Test Box Assembly (Part No. 9870005-1) is
available from Cessna Parts Distribution (CPD 2),
through Cessna Service Stations, for use in isolating
failures in the 28-volt transistorized voltage regulator
(C611002-0105) and the 28-volt alternator.
17-42. REMOVAL AND INSTALLATION. (Refer to
figure 17-5).
a. Ensure that the master switch is off.
b. Remove upper cowl to gain access to the regulator.
c. Remove the connector plug from the regulator.
d. Remove the three bolts holding the regulator on
the firewall.
e. To reinstall the regulator, reverse the preceding
steps.
17-42A. ALTERNATOR CONTROL UNIT.
NING WITH 1979 MODELS.)
(BEGIN-
17-42B. DESCRIPTION. The alternator control unit
is a solid state voltage regulator with an over-voltage
sensor and a low-voltage sensor incorporated in the unit.
The control unit is not adjustable and is a remove-andreplace item. A Cessna Alternator Charging System
Test Box Assembly (Part No. 9870005-1) is available
from Cessna Parts Distribution (CPD 2), through Cessna
Service Stations, for use in isolating failures in the
28-volt alternator control units (C611005-0101 and
C611005-0102) and the 28-volt alternator.
17-42C. REMOVAL AND INSTALLATION. (Refer to
figure 17-5.)
a. Thru 1980 Models remove upper half of engine
cowl. Beginning with 1981 Models the control unit
is mounted on the aft side of the battery box, under
the instrument panel.
b. Place master switch in the "OFF"position.
c. Disconnect negative lead from the battery.
d. Disconnect housing plug from the alternator
control unit.
e. Remove screws securing the control unit to the
firewall.
f. To install control unit reverse the preceding
steps. Be sure the connections for grounding are
clean and bright before assembly. Otherwise faulty
voltage regulation and/or excessive radio noise may
result.
17-43. OVER-VOLTAGE SENSOR AND WARNING
LIGHT.
17-30
Revision3
17-44. DESCRIPTION. The over-voltage system
consists of a over-voltage sensor switch and a red
warning light labeled, HIGH VOLTAGE, on the instrument panel. When an over-voltage tripoff occurs the
over-voltage sensor turns off the alternator system
and the red warning light comes on. The ammeter
will show a discharge. Turn off the alternator portion
of the master switch to recycle the over-voltage sensor. If the over-voltage condition was transient, the
normal action is necessary. If the over-voltage tripoff recurs, then a generating system malfunction has
occurred such that the electrical accessories must be
operated from the aircraft battery only. Conservation
of electrical energy must be practiced until the flight
can be terminated. The over-voltage red warning
light filament may be tested at any time by turning off
the alternator portion of the master switch and leaving
the battery portion turned on. This test does not induce an over-voltage condition on the electrical system. Beginning with 1979 Models the over-voltage
sensor is contained within the alternator control unit.
The unit also contains a low-voltage sensor. A red
warning light labeled "LOW VOLTAGE" is installed
on the instrument panel. When an over-voltage condition occurs the over-voltage sensor turns off the
alternator and the voltage in the system drops. When
system voltage drops below 24.8 volts the low-voltage
sensor turns on the low-voltage light indicating a
drain on the battery and the ammeter will show a discharge. Turn off both sections of the master switch
to recycle the over-voltage sensor. If the overvoltage condition was transient, the normal alternator
charging will resume and no further action is necessary. If the over-voltage tripoff recurs, then a generating system malfunction has occurred such that
the electrical accessories must be operated from the
aircraft battery only. Conservation of electrical
energy must be practiced until the flight can be terminated. The over-voltage light filament may be
tested at any time by turning off the "Alternator" portion of the master switch and leaving the battery portion on. This test does not induce an over-voltage
condition on the electrical system.
NOTE
On 1979 thru 1982 models if the alternator low
voltage light comes on when a COM radio transmitter is keyed, refer to Cessna Single Engine
Customer Care Service Information Letter SE8217 Dated April 30, 1982.
17-45. REMOVAL AND INSTALLATION. (Refer to
figure 17-6. )
a. Turn master switch (BATT side) to OFF position.
b. Disconnect plug.
c. Remove mounting screws and remove relay.
d. To install reverse the procedure.
17-46. RIGGING THROTTLE-OPERATED MICROSWITCH. Refer to Section 13.
17-47. AUXILIARY FUEL PUMP FLOW RATE
ADJUSTMENT. Refer to Section 13.
MODEL 210 & T210 SERIES SERVICE MANUAL
4 THRU 1978 MODELS
21063473
1.
2.
3.
4.
5.
6.
7.
8.
Housing - Cap
13
Wire (to Alternator Ground)
Voltage Regulator
Screw
Housing - Plug
Cover
Sta-strap
Clamp
1979 THRU 1980 MODELS
Detail A
9.
10.
11.
12.
13.
14.
15.
16.
Alternator Control Unit
Terminal Block
Spiral Wrap
Wire (to Circuit Breaker)
Wire (to Alternator Control Unit)
Wire (to Alternator)
Ground Wire
Spacer
Figure 17-5. Voltage Regulator/Alternator Control Unit Installation (Sheet 1 of 2)
17-31
MODEL 210 & T210 SERIES SERVICE MANUAL
2
3
8
Detail A
1.
2.
3.
4.
5.
6.
7.
8.
Sta-strap
Housing Cap
Housing Plug
Alternator Control Unit
Bracket
Battery Box
Ground Wire
Bolt
BEGINNING WITH 1981 MODELS
Figure 17-5.
17-32
Voltage Regulator/Alternator Control Unit Installation (Sheet 2 of 2)
MODEL 210 & T210 SERIES SERVICE MANUAL
17-47A.
STANDBY GENERATOR SYSTEM.
17-471.
ALTERNATOR CONTROL UNITS.
17-47B. DESCRIPTION. The standby generator
system may be installed on the aircraft beginning
with 1980 models. The system provides a 24 volt DC,
7-amp capacity of standby power for the following
essential electrical and avionic equipment in the event
event the main electrical system cannot be used; gear
warning, stall warning, fuel quantity, turn coordinator,
engine oil and cylinder head temp, also circuit breaker
(radio 3) and (radio 1 or 2). The system consists of a
standby generator, mounted on the engine accessory
case. a voltage regulator, mounted on the upper right
hand portion of the firewall, a two-position toggle OFFON switch and a two-position toggle radio selector
switch (labeled NC1/NC2) installed on the circuit breaker panel. For trouble shooting and adjustments refer
to the Standby Generator Charging Systems Manual,
D5021-13, dated 15 September 1979.
17-47J. DESCRIPTION. The alternator control units
are solid state voltage regulators with low voltage
sensing internal paralleling circuitry in the alternator
control units controls load sharing between the alternators.
17-47C. REMOVAL AND INSTALLATION. Refer to
figure 17-6A.
electrcal system operation.
17-47N. REMOVALAND INSTALLATION.
17-47D.
DUAL ALTERNATOR SYSTEM.
ALTERNATORS.
17-47G. DESCRIPTION. The alternators are beltdriven, 28 volt, 60 amp, three-phase, Delta connected stator windings with integral silicon diode
rectifiers and a stator tap.
NOTE
Alternators are equal in function & capability,
and normally operate under equal loads. Each
may operate independently, but should not be
thought of or operated as, a primary and
secondary (or standby) system.
17-47H. REMOVAL AND INSTALLATION.
figure 17-6B.)
17-47L.
(See
ALTERNATOR CONTACTORS AND SHUNTS
17-47M. DESCRIPTION. Each alternator is equipped
with a contactor and shunt. The shunt directs power
through two fuses to the alternator control unit remote
sensing and current sensing circuits. The shunt is
also connected through fuses to the volt-ammeter
selector switch which enables the pilot to monitor the
(See
figure 17-6B.)
17-47E. DESCRIPTION. The dual alternator system
consists of two belt-driven, 28 volt, 60 amp alternators, two alternator control units, two shunt and fuse
assemblies, two line contactors, two alternator
switches, two circuit breakers, a volt ammeter, a
three light indicating system and a alternator restart
system. An isolation circuit breaker is installed with
the dual alternator system. Refer to the Pilots Operating Handbook for operational procedures.
17-47F.
17-47K. REMOVAL AND INSTALLATION.
figure 17-6B.)
(See
17-470.
VOLT-AMMETER.
17-47P. DESCRIPTION.The volt-ammeter is
mounted on the left side of the nstrument panel. A
selector switch is provided for the pilot to monitor
the electrical system operation. The selector switch
allows the pilot to monitor the current supplied by
each alternator, the battery charge or discharge current, or the system voltage.
17-47Q. ALTERNATOR RESTART SYSTEM. The
alternator restart system consists of a battery pack
and a switch. When the restart switch, on the circuit
breaker panel is actuated, power is directed from the
battery pack through the restart switch to the alternator switch. With the alternator switch closed power
is directed to the alternator control unit then to the
alternator field for excitation of the alternator.
NOTE
Batteries should be changed at yearly Intervals or sooner if function test shows need.
Correct polarity must be observed when
installing batteries. No. 814 Ray-O-Vac
or No. MN1400 Mallory or equivalent to
No. E-93 Everready Batteries are recommended.
WARNING
Do not rely on contact between battery holder
(78) and plate (79) to maintain spring contact
on batteries. If required, end plates of the
battery holder may be reformed inward slightly to increase contact pressure on batteries.
Check continuity of battery pack before installation with battery pack suspended from plate
and with curvature of plate reversed as in
normal installation.
17-34
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
•TO RQUE 450 - 500 IN-LBS
*TORQUE 160 - 190 IN-LBS
*TORQUE 155 - 175 IN-LBS
36
21.Mount
34
\
t
23
26. Bolt
4%
27. Adjustment Bracket
28. Nipple
B 21064570
Detail
WITH
BlINNING
DetilB
34
BEGINNING WITH
~33 21064570
Shunt)
29. Wire
Wire (Alternator
31.
(Ground) Sense)
32. Wire (Remote
33. Bolt
34. Wire (Alt OFF Sense)
35. Resistor
Shunt)
Wire
Alternator
4029.
Insulator
36.
Wire Sense)
Safety(Remote
Resistor
37. ~25~38.
Wire
31.
Washer
a3041.
Bolt
26.
32.
(Ground)
Bracket
Adjustment
27, Wire
33. Bolt
2839.NipplScrew
Figure17-6B34.
Installation
AlternatorFF
(Sheet
Sytem4 WireDual
of 8)
36. Insulator Resistor37.
38.
39.
40.
41.
Figure 17-6B.
1735.
Resistor
Washer
Screw
Alternator
Safety Wire
Dual Alternator System Installation (Sheet 4 of 8)
17-39
MODEL 210 & T210 SERIES SERVICE MANUAL
41.
42.
43.
44.
Cover
Wire (Circuit Breaker)
Contactor
Bracket (Alternator No. 2)
45. Bus Bar
46. Shunt (Alternator No. 2)
47. Shunt (Alternator No. I)
48. Wire (to Alternator No. 1)
49. Wire (to Alternator No. 2)
50. Wire (Circuit Breaker)
Figure 17-6B. Dual Alternator System Installation (Sheet 5 of 8)
17-40
MODEL 210 & T210 SERIES SERVICE MANUAL
52
51
54
55
DetailD
56
60
59
61
62
51. Cover
52. Shunt (Battery)
53. Sleeve
54. Wire (to Main Bus)
55. Wire (to Battery Contactor)
56. Tie
57. Strap
58.
59.
60.
61.
62.
Insulator
Diodes
Bus Bar
Bus Bar (Dual Alternator only)
Isolation Circuit Breaker
Detail l
Figure 17-6B. Dual Alternator System Installation (Sheet 6 of 8)
17-41
MODEL 210 & T210 SERIES SERVICE MANUAL
17-48. AIRCRAFT LIGHTING SYSTEM.
17-50. SWITCHES.
17-49. DESCRIPTION.
The aircraft lighting systern consists of landing and taxi lights, navigation
lights, flashing beacon light, anti-collision strobe
lights, interior and instrument panel flood lights,
electroluminescent panel lighting, instrument post
lighting, pedestal lights, oxygen lights, courtesy
lights, de-ice light, control wheel map light, baggage compartment light, compass and radio dial
lights.
17-51. DESCRIPTION. The instrument panel
switches used are snap-in type rocker switches.
These switches have a design feature which permits
them to snap into the panel from the panel side and
can subsequently be removed for easy maintenance.
These switches also feature spade type slip-on terminals.
17-52. TROUBLE SHOOTING.
TROUBLE
LANDING AND TAXI LIGHTS
OUT.
LANDING OR TAXI LIGHT
OUT.
FLASHING BEACON DOES
NOT LIGHT.
FLASHING BEACON
CONSTANTLY LIT.
17-44
PROBABLE CAUSE
Short circuit in wiring.
REMEDY
1. Inspect circuit breaker. If
circuit breaker is open, proceed
to step 2. If circuit breaker is
OK, proceed to step 3.
Defective wiring.
2. Test each circuit separately
until short is located. Repair
or replace wiring.
Defective switch.
3. Check voltage at lights with
master and landing and taxi light
switches ON. Should read battery voltage. Replace switch.
Lamp burned out.
1. Test lamp with ohmmeter or
new lamp. Replace lamp.
Open circuit in wiring.
2. Test wiring for continuity.
Repair or replace wiring.
Short circuit in wiring,
1. Inspect circuit breaker. If
circuit breaker is open, proceed
to step 2. If circuit breaker is
OK, proceed to step 3.
Defective wiring.
2. Test circuit until short is located. Repair or replace wiring.
Lamp burned out.
3. Test lamp with ohmmeter or
a new lamp. Replace lamp. If
lamp is good, proceed to step 4.
Open circuit in wiring.
4. Test circuit from lamp to
flasher for continuity. If no
continuity is present, repair or
replace wiring. If continuity is
present, proceed to step 5.
Defective switch.
5. Check voltage at flasher with
master and beacon switch on.
Should read battery voltage.
Replace switch. If voltage is
present. proceed to step 6.
Defective flasher.
.
Defective flasher.
1.
Install
flasher.
Install new flasher.
MODEL 210 & T210 SERIES SERVICE MANUAL
17-52. TROUBLE SHOOTING (Cont.)
TROUBLE
PROBABLE CAUSE
ALL NAV LIGHTS OUT.
ONE NAV LIGHT OUT.
Short circuit in wiring.
REMEDY
1. Inspect circuit breaker. If
circuit breaker is open, proceed
to step 2. If circuit breaker is
OK, proceed to step 3.
Defective wiring.
2. Isolate and test each nav light
circuit until short is located.
Repair or replace wiring.
Defective switch.
3. Check voltage at nav light with
master and nav light switches on.
Should read battery voltage. Replace switch.
Lamp burned out.
1. Inspect lamp.
Open circuit in wiring;
2. Test wiring for continuity.
Repair or replace wiring.
Replace lamp.
WARNING
The anti-collision system is a high voltage device. Do not remove
or toach tube assembly while in operation. Wait at least 5 minutes
after turning off power before starting work.
BOTH ANTI-COLLISION
STROBE LIGHTS WILL
NOT LIGHT.
Open circuit breaker.
1. Check, if open reset. If
circuit breaker continues to
open proceed to step 2.
2. Disconnect red wire between aircraft power supply
(battery/external power) and
strobe power supplies, one
at a time. If circuit breaker
opens on one strobe power
supply, replace strobe power
supply. If circuit breaker
opens on both strobe power
supplies proceed to step 3.
If circuit breaker does not
open proceed to step 4.
3. Check aircraft wiring.
Repair or replace as necessary.
4. Inspect strobe power supply ground wire for contact
with wing structure.
17-45
MODEL 210 & T210 SERIES SERVICE MANUAL
17-52. TROUBLE SHOOTING (Cont.)
TROUBLE
PROBABLE CAUSE
REMEDY
CAUTION
Extreme care should be taken when exchanging flash tube. The tube
is fragile and can easily be cracked in a place where it will not be
obvious visually. Make sure the tube is seated properly on the base
of the nav light assembly and is centered in the dome.
NOTE
When checking defective power supply and flash tube, units from
opposite wing may be used. Be sure power leads are protected
properly when unit is removed to prevent short circuit.
ONE ANTI-COLLISION
STROBE LIGHT WILL
Defective Strobe Power Supply,
or flash tube.
1. Connect voltmeter to red lead
between aircraft power supply
(battery/external power) and
strobe power supply, connecting
negative lead towing structure.
'Check for 12/24 volts. If OK proceed to step 2. I not, check aircraft power supply (battery/external power).
2. .Replace flash tube with known
good flash tube. If system still
does not work, replace strobe
power supply.
DOME LIGHT TROUBLE.
Short circuit in wiring.
Defective wiring.
1. Inspect circuit breaker. If
circuit breaker is open, proceed
to step 2. If circuit breaker is
OK, proceed to step 3.
2. Test circuit until short is
located. Repair or replace
wiring.
3. Test for open circuit. Repair
or replace wiring. If no short or
open circuit is found, proceed to
step 4.
17-46
Lamp burned out.
4. Test lamp with ohmmeter or
new lamp. Replace lamp.
Defective switch.
5. Check for voltage at dome
light with master and dome light
switch on. Should read battery
voltage. Replace switch.
MODEL 210 & T210 SERIES SERVICE MANUAL
17-52. TROUBLE SHOOTING (Cont.)
TROUBLE
ELECTROLUMINESCENT
PANELS WILL NOT LIGHT.
PROBABLE CAUSE
REMEDY
Short circuit in wiring.
1. Inspect circuit breaker. If
circuit breaker is open, proceed
to step 2. If circuit breaker is
OK, proceed to step 3.
Defective wiring.
2. Test circuit until short is
located. Repair or replace wiring.
3. Test for open circuit. Repair
or replace wiring. If no open or
short circuit is found, proceed to
step 4.
Defective resistor.
4. Check resistor for continuity.
(Located in line between rheostat
and inverta-pak.) Replace resistor.
Defective rheostat.
5. Check input voltage at invertapak with master switch on. Voltmeter should give a smoothly varied
reading over the entire control range
of the rheostat. If no voltage is present or voltage has a sudden drop
before rheostat has been turned full
counterclockwise, replace rheostat.
Defective inverta-pak.
6. Check output voltage at invertapak with ac voltmeter. Should read
about 125 volts ac with rheostat set
for full bright. Replace inverta-
pak.
INSTRUMENT LIGHTS WILL
NOT LIGHT,
Short circuit wiring.
1. Inspect circuit breaker. If
circuit breaker is open, proceed to
step 2. If circuit breaker is OK,
proceed to step 3.
Defective wiring.
2. Test circuit until short is located. Repair or replace wiring.
3. Test for open circuit. Repair
or replace wiring. If no short or
open circuit is found, proceed to
step 4.
Faulty section in
dimming potentiometer.
4. Lights will work when control
is placed in brighter position. Replace potentiometer.
Faulty light dimming
transistor.
5. Test both transistors with new
transistor. Replace faulty transistor.
Faulty selector switch.
6.
Inspect.
Replace switch.
17-47
MODEL 210 & T210 SERIES SERVICE MANUAL
17-52. TROUBLE SHOOTING (Cont.)
PROBABLE CAUSE
TROUBLE
INSTRUMENT LIGHTS WILL
NOT DIM.
REMEDY
Open resistor or wiring
in minimum intensity end
of potentiometer.
1. Test for continuity. Replace
resistor or repair wiring.
Shorted transistor.
2. Test transistor by substitution.
Replace defective transistor.
Nav light switch turned off.
1. Nav light switch has to be
ON before map light will light.
Short circuit in wiring.
2. Check lamp fuse on terminal
board located on back of stationary panel with ohmmeter. If
fuse is open, proceed to step 3.
If fuse is OK, proceed to step 4.
Defective wiring.
3. Test circuit until short is located. Repair or replace wiring.
CONTROL WHEEL MAP
LIGHT WILL NOT LIGHT.
4. Test for open circuit. Repair
or replace wiring. If a short or
open circuit is not found, proceed
to step 5.
Defective map light assembly.
17-53. LANDING AND TAXI LIGHTS.
17-54. DESCRIPTION. The landing and taxi lights
are mounted in the lower nose cap. Both lamps are
used for landing and only the right hand for taxi thru
1977 models and the left beginning with 1978 models.
The lamps are controlled by two rocker switches
with a diode assembly installed across the switches
which enable the landing light switch to turn on both
the landing and taxi lamps. The taxi light switch will
turn on only the taxi lamp.
17-55. REMOVAL AND INSTALLATION. (Refer to
figure 17-7.)
a. Remove screws securing retainer (2) to nose
cap.
b. Pull light assembly forward from nose cap and
disconnect lamp wires.
c. Remove tinnerman screws (6) from bracket (5)
and remove bracket and lamp.
d. Install new lamp and reassemble.
17-56. NAVIGATION LIGHTS.
17-57. DESCRIPTION. The navigation lights are
located on each wing tip and the stinger. Operation
of the lights is controlled by a single two position
switch. A plastic light detector on each wing tip
allows the pilot to determine if the lamps are working properly during flight.
17-48
5. Check voltage at map light
assembly with master and nav
switches on. If battery voltage
is present, replace map light
assembly.
17-58. REMOVAL AND INSTALLATION. Refer to
figure 17-8 for removal and installation of navigation
light components.
17-59. ANTI-COLLISON STROBE LIGHTS.
17-60. DESCRIPTION. A white strobe light may be
installed on each wing tip with the navigation light.
These lights are vibration resistant and operate on
the principle of a capacitor discharge into a zenon
tube, producing an extremely high intensity flash.
Each strobe light has its own power supply mounted
on the wing tip ribs.
17-61. OPERATIONAL REQUIREMENTS.
(THRU 1977 MODELS).
WARNING
The capacitors in the strobe light power
supplies must be reformed if not used for
a period of six (6) months. The following
procedure must be used.
Connect the power supply, red wire to plug, black to
ground to 6 volt DC source. Do Not connect strobe
tube. Turn on 6 volt supply. Note current draw after
one minute. If less than 1 ampere, continue operation for 24 hours. Turn off DC power source. Then
connect to the proper voltage, 24 volt. Connect tube
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
to output of strobe power supply and allow to operate,
flashing, for 15 minutes. Remove strobe tube. Operating power supply at 24 volts, note the current
drain after one minute. If less than 0. 5 amperes,
operate for 6 hours. If current draw is greater than
0. 5 amperes, reject the unit.
WARNING
This anti-collision system is a high voltage
device. Do not remove or touch tube assembly while in operation. Wait at least 5 minutes after turning off power before starting
work.
17-62. REMOVAL AND INSTALLATION. Refer to
figure 17-8 for removal and installation of strobe
light components.
a. Remove wing tip disconnecting navigation and
strobe light wires.
b. Disconnect power supply wires.
c. Remove the four mounting screws and remove
power supply.
d. To reinstall reverse the preceding steps.
17-62A. VERTICAL TAIL FLOOD LGHTS.
17-62B. DESCRIPTION. A flood light assembly is
mounted on each end of the stabilizer, on the upper
side. These lights are used to illuminate the vertic-
al tail. A switch on the switch panel controls the
lights and a circuit breaker on the breaker panel
protects the circuit.
17-62C. REMOVAL AND INSTALLATION. Refer
to figure 17-8. for removal and installation.
NOTE
To properly secure the lens (4) to the
fixture, 5 in-lbs (min) to 6 in-lbs (max)
should be used. The screw should be
tightened to the point that the lens is
properly seated on the gasket and the
"O" ring under the hold down screw
washer is compressed without undue
strain on the glass.
17-50
Revision 2
NOTE
Aircraft equipped with light assemblies
using either 28 volt lamps or 14 volt
lamps connected in series. 14 volt lamps
assemblies are identified by rubber
stamping "14V" on the lamp base. Refer
to applicable wiring diagram if in doubt.
It is imperative that 14 volt lamps are
not installed in the 28 volt light assemblies as this will result in the immediate
burn out of the lamp. Should 28 volt lamps
be installed in the 14 volt light assemblies, there will be a considerable reduction of light output.
17-63. FLASHING BEACON
17-64. DESCRIPTION. The flashing beacon light is
attached to the vertical fin tip. The flashing beacon
has a iodine-vapor lamp electrically switched by a
solid-state flasher assembly. The flasher assembly
is mounted inside the fin tip. The switching frequency
of the flasher assembly operates at approximately 45
flashes per minute. A resistor is installed and connected to the unused flasher lead to eliminate a pulsing effect on the cabin lighting and ammeter.
17-65. REMOVAL AND INSTALLATION. Refer to
figure 17-9 for removal and installation of flashing
beacon components.
17-66. INSTRUMENT LIGHTING.
17-67. DESCRIPTION. The instrument panel lighting consists of two seperate sections. The lower
two-thirds of the panel is illuminated by two lights
mounted in the overhead console. The lighting for
the upper one-third of the panel is provided by four
lights mounted in the under side of the instrument
glare shield. The Intensity of the lighting is controled by the instrument light dimming rheostat located
on the switch panel.
17-68. REMOVAL AND INSTALLATION. Refer to
figure 17-10 for removal and installation of instrument
brow lights.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
B
C
CA
1. Housing
Figure 17-10. Instrument Panel Glare Shield Light Installation
17-55
MODEL 210 & T210 SERIES SERVICE MANUAL
17-69. REMOVAL AND INSTALLATION OF OVERHEAD CONSOLE INSTRUMENT PANEL LIGHTS.
(Refer to figure 17-11).
a. Unscrew metal oxzgen port covers, if installed.
b. Unscrew oxygen gage lens, if installed.
c. Remove screw from oxygen control knob and
remove knob.
d. Remove the screws in the recess area of the
fresh air vents.
e. Pull out the two oxygen post lights, if installed.
f. Remove remaining screws the over-head console
cover and remove cover.
g. Twist lamp for removal from socket assembly.
h. For installation, reverse the preceeding steps.
17-70. VERTICAL ADJUSTMENT OF OVERHEAD
CONSOLE INSTRUMENT PANEL LIGHTS. (Refer
to figure 17-11).
a. Pry the plug button from the overhead console
cover to gain access to the adjustment screw.
b. Turn the screw clockwise to advance the light
beam up the panel.
c. Turn the screw counterclockwise to advance the
light down the panel.
d. Upon completing adjustment, reinstall plug
button.
17-71. LATERAL ADJUSTMENT OF OVERHEAD
CONSOLE INSTRUMENT PANEL LIGHTS. (Refer
to figure 17-11).
a. To gain access to the lights, remove the overhead console cover as outlined in paragraph 17-69.
b. Slide the light sockets inboard along the mounting bracket to advance the light beam outboard on the
instrument panel. To advance the light beam inboard
on the instrument, slide the light socket outboard
along the mounting bracket.
consists of a two-circuit transistorized dimming
assembly, mounted on the right hand side of the cabin
forward of the instrument panel, and two controls on
the lower left hand side of the panel. The left control
is a dual rheostat with a concentric knob arrangement.
The center portion controls lower panel lighting, the
outer portion controls engine instrument and radio
lighting. The right hand control is a single rheostat
and controls instrument lighting. This includes,
glare shield lights, instrument flood lights, compass
light and post lighting if installed. Beginning with
1978 Models a three-circuit transistorized dimming
assembly is installed with post lighting.. The controls
go from three to four with the post light installation.
The center portion of the left hand control, controls
the post lights, the outer portion controls flood lights,
the center portion of the right hand control, controls
E L panel lighting and the outer portion controls
engine and radio lighting.
17-76. REMOVAL AND INSTALLATION
For removal and installation of transistorized dimming,
refer to figure 17-12.
17-77. PEDESTAL LIGHTS.
17-78. DESCRIPTION. The pedestal lights consist
of three post type lights mounted on the pedestal to
illuminate the fuel selector handle, rudder and elevator trim controls. The pedestal lights are controlled by the instrument light rheostat.
17-79. REMOVAL AND INSTALLATION. For removal and installation of pedestal lamps, slide the
cap and lens assembly from the base. Slide the lamp
from the socket and replace.
17-80. INSTRUMENT POST LIGHTING.
NOTE
Should sliding the light sockets along the mounting bracket prove difficult, the screws attaching the light socket assembly to the mounting
bracket may be loosened to permit the light
socket assembly to slide along the mounting
bracket. Once the adjustment is completed,
ensure that the screws are tight enough to
resist vibrating out of adjustment.
17-72. ELECTROLUMINESCENT PANEL LIGHTING.
17-73. DESCRIPTION. The electroluminescent
lighting consists of two "EL" panels; the switch panel
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. These "EL" panels have an expected life of
over 16, 000 hours and no replacement should be
necessary during the life of the aircraft.
17-74. TRANSISTORIZED LIGHT DIMMING.
17-75. DESCRIPTION.
17-56
The light dimming circuit
17-81. 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 instrument light dimming rheostat located
on the switch panel.
17-82. 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.
17-83. OXYGEN LIGHTS.
17-84. DESCRIPTION. The oxygen lights consist of
two post type lights installed in the overhead oxygen
console. The intensity of the oxygen lights is controlled by the radio light dimming rheostat located
on the switch panel.
17-85. REMOVAL AND INSTALLATION. Refer to
figure 17-11 and paragraph 17-82 for removal and inst
installation of oxygen post lights.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
17-86. COURTESY LIGHTS.
17-87. 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.
17-91. REMOVAL AND INSTALLATION. (Refer to
figure 17-16.)
a. Ensure that the master switch is "OFF".
b. To gain access to the baggage compartment
lamp, remove the screws attaching the retainer
and lens to the reflector assembly.
c. Twist the lamp from the socket.
d. To replace the bulb, reverse this procedure.
17-88. REMOVAL AND INSTALLATION. Refer to
17-92. INTERIOR LIGHTING
figure 17-13 for removal and installation of courtesy
lights.
17-93..
DESCRIPTION. Interior lighting consists of
a dome light installed in the overhead console aft of
17-89. BAGGAGE COMPARTMENT LIGHT.
rear wing spar. A slide switch located forward of
17-90. DESCRIPTION. The baggage compartment
is illuminated by a lamp mounted in the top of the
baggage compartment. The light is controlled by the
"Utility Lights" switch located on the left door post.
the light controls the lamp.
17-94. REMOVAL AND INSTALLATION.
a. Snap lens out of cover.
b. Remove lamp and replace with new lamp.
c. Reinstall lens.
3
4
2
A
1. Grommet
2. Screw
3. Shield
4. Socket
5. Lamp
DetailA
Figure 17-13.
6. Cover Plate
7. Tinnerman Nut
8. Spacer
9. Lens Assembly
10. Cover Assembly
Courtesy Light Installation
17-59
MODEL 210 & T210 SERIES SERVICE MANUAL
8
9 10
10
1
2
14
7
THRU 1977 MODELS
2
13
12 14
13
BEGINNING WITH 1978 MODELS
1.
2.
3.
4.
5.
6.
7.
Control Tube Assembly
Cover
Adapter
Connector
Plate
Map Light Rheostat
Control Wheel
8.
9.
10.
11.
12.
13.
14.
Pad
Mike Switch
Plug
Insulator
Map Light Assembly
Lamp
Knob (Map Light)
Figure 17-14. Control Wheel Map Light Installation
17-60
MODEL 210 & T210 SERIES SERVICE MANUAL
Deleted
17-95. CONTROL WHEEL MAP LIGHT.
17-100.
17-96. DESCRIPTION. The control wheel mwp light
is internally mounted in the control wheel. A rheostat on the lower left hand side of the wheel controls
17-101. Deleted
the light.
17-102. Deleted
17-103.
17-97. REMOVAL AND INSTALLATION. (Refer to
figure 17-14.) To remove lamp. push upward on the
lamp and turn. The lamp and reflector are replaced
as a unit.
17-104. DESCRIPTION. A solid state warning unit is
installed on the right hand wing root rib. The warning
siginal is transmitted through the radio speaker in the
overhead console.
17-98. COMPASS AND RADIO DIAL LIGHTS.
17-99. DESCRIPTION. The compass and radio dial
lights are contained within the individual units. The
light intensity is controlled by the instrument light
dimming rheostat mounted on the lower left side of
the instrument panel.
STALL WARNING UNIT.
NOTE
On Aircraft Serials 21061040 thru 21062249
i false signals are experienced Refer to
Cessna Single-engine Service Letter SE7850 dated August 7, 1978.
Figure 17-15. Deleted
17-61
MODEL 210 & T210 SERIES SERVICE MANUAL
_./
1. Screw
2. Grommet
3. Sta. 138 Bulkhead
4. Bracket
5.
6.
7.
Figure 17-16.
17-62
Nutplate
Reflector
Nut
Baggage Compartment Light Installation
8.
9.
10.
11.
Retainer
Lens
Lamp
Socket
MODEL 210 & T210 SERIES SERVICE MANUAL
17-105. REMOVAL AND INSTALLATION. Refer to
figure 17-17 for removal and installation.
17-109. PITOT AND STALL WARNING HEATERS.
17-110. DESCRIPTION. Electrical heater units are
incorporated in some pitot tubes and stall warning
switch units. The heaters offset the possibility of
ice formation on the pitot tube and stall warning actuator switch. The heaters are integrally mounted in
the pitot tube and stall warning actuator switch.
Both heaters are controlled by the pitot heat switch.
17-106. STALL WARNING SWITCH.
17-107. DESCRIPTION. The stall warning 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 stall warning horn. The horn should
sound at approximately five to ten miles per hour
above the actual stall speed. Initial installation of
the switch should be with the Up of the warning switch
approximately one sixteenth of an inch below the center line of the wing skin cutout. Test fly the aircraft
to determine if the horn sounds at the desired speed.
If the horn sounds too soon, move the unit down
slightly; if too late, move the unit up slightly.
17-111. REMOVAL AND INSTALLATION Refer to
figures 17-17 and 17-18 for removal and installation.
17-112. LANDING GEAR INDICATOR LIGHTS.
17-113. DESCRIPTION. The position of the landing
gear is indicated by two press-to-test lamp assemblies mounted on the right side of the switch panel.
The green light is on when all the wheels are down
and locked; the amber is on when all the wheels are
17-108. REMOVAL AND INSTALLATION. Refer to
figure 17-17 for removal and installation.
A
b
M.
:1__
1.
2.
3.
4.
Dual Warning Unit
Adjustment Pots
RH Wing Root Rib
Screw
5. Cover
Figure 17-17.
Stall Warning Unit
17-63
MODEL 210 & T210 SERIES SERVICE MANUAL
up and locked. If any wheel assumes an intermediate
position of neither up and locked or down and locked,
both lights will be dark. The hood of each Light is
removable for bulb replacement, and has a dimming
shutter.
17-114. REMOVAL AND INSTALLATION.
a. Remove the hood on either light by unscrewing
counterclockwise. The lamp bulb is in the hood and
may be replaced by pulling it out and inserting a new
lamp.
b. To remove the lamp socket assembly, remove
the nut from the assembly on the front side of the
panel.
c. Tag and unsolder the wires from the socket
assembly.
d. To replace a lamp socket assembly, reverse
the above procedure.
17-115. LANDING GEAR WARNING HORN.
Refer to Section 5.
17-116. CIGAR LIGHTER. (THRU 21064536)
17-117. DESCRIPTION. A special circuit breaker
is 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.
CAUTION
Make sure the masterswitchis "OFF"
before inserting probe into the-circuit
breaker on cigar lighter to reset.
17-118. REMOVAL AND INSTALLATION. (Refer to
figure 17-20).
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.
A
1. Wing Skin
2.
3.
4.
Actuator
Tinnerman Nut
Screw
4
Figure 17-18.
SHOP NOTES:
17-64
Stall Warning Switch.
Detail A
MODEL 210 & T210 SERIES SERVICE MANUAL
1. Electrical Leads
2. Pitot Tube
3. Heating Element
DetailA
Figure 17-19. Pitot Heater
THRU 21064536
Figure 17-20.
1.
2.
3.
4.
5.
Knob
Element
Socket
Panel
Shell
6.
Circuit Breaker
7.
8.
9.
10.
Probe
Nut
Lockwasher
Power Wire
Cigar Lighter Installation
17-65
MODEL 210 &T210 SERIES SERVICE MANUAL
17-119.
Deleted.
CAUTION
17-120.
Deleted.
17-121.
Deleted.
17-122.
Deleted.
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.
17-123. EMERGENCY LOCATOR TRANSMITTER.
THRU 21061715.
17-124. 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 reception by search 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 are equipped with a battery pack containing four lithium "D" size batteries which are stacked
in two's (See figure 17-23). 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
17-126. OPERATIONAL TEST OF EMERGENCY
LOCATOR SYSTEM. The ELT, its battery pack, and its
antenna must be inspected and tested each 100 hours.
The operational test of the airplane's emergency locator
system should check both radiated signal strength and
the ELT G-switch. The airplane's VHF receiver is
located very close to the ELT and is very sensitive.
Consequently, using the airplane's VHF receiver to
monitor ELT transmission does not provide same level of
confidence in verifying ELT signal as using AM radio or
performing control tower check.
CAUTION
Tests with the antenna connected should
be approved by the nearest control
tower. The FAA/DOT allows free space
transmission tests from the airplane only
within first five minutes after each hour.
The test time allowed is limited to three
sweeps of the warble tone or
approximately one second control
tower should be notified that a test is
about to be conducted.
NOTE
NOTE
After accumulated test or operation time
in the following table.
equals one hour, battery pack
replacement is required.
TRANSMITTER LIFE
TO 75 MILLIWATTS OUTPUT
Temperature
4-Cell
Lithium
Battery Pack
115
115
95
23
hrs
hrs
hrs
hrs
Battery packs have a normal shelf life of five to ten
(5-10) years and must be replaced at half of normal shelf
life in accordance with TSO-C91. Cessna specifies 5
years replacement of lithium (4-cell) battery packs.
17-125. 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 -0g) impact force.
for a duration of 11-16 milliseconds.
17-66
Revision 3
Operational test of radiated signal with control
(1) Turn airplane master switch ON.
(2) Verify that test is conducted within first five
__~
_____________
+130*F
- 70°F
- 4F
- 40°F
a.
tower monitoring.
minutes of the hour.
(3) Turn airplane transceiver ON, request
permission from nearest control tower and flight service
station to conduct operational test of ELT, and request
control tower monitoring.
(4) Place ELT function selector to the ON position
for one second or less (no more than three sweeps of the
audio signal). Immediately replace the ELT function
selector to the ARM position after testing ELT.
(5) Contact control tower and confirm proper
locator beacon operation.
(6) Restore switches to normal.
b. Operational test of radiated signal with handheld
AM radio monitoring.
(1) Turn airplane master switch ON.
(2) Verify that test is conducted within first five
minutes of the hour.
(3) Turn airplane transceiver ON and request
permission from nearest control tower and flight service
station to conduct operational test of ELT.
(4) Position a small hand held AM radio tuned to
any frequency within six inches of the ELT antenna.
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
(5) Place ELT function selector to the ON position
for one second or less (no more than three sweeps of the
audio signal). Immediately replace the ELT function
selector to the ARM position after testing ELT.
(6) Verify that ELT signal has been detected on
hand held AM radio.
(7) Restore switches to normal.
c. Operational test of the TSO-C91 ELT G-switch.
(1) Remove ELT from airplane.
(2) While holding ELT in one hand, sharply strike
the end of the case in the direction of activation
indicated on the case of the transmitter.
(3) Using either radiated signal test method
described above, verify that the G-switch has been
activated and ELT is transmitting.
(4) Reset the G-switch, and restore other
disturbed switches to normal.
(5) Reinstall ELT in airplane.
d. Operational test of the TSO-C9la ELT G-switch.
(1) Remove ELT from airplane.
(2) While holding ELT firmly in one hand, make a
throwing motion followed by a sudden reversal of the
transmitter.
(3) Using either radiated signal test method
described above, verify that the G-switch has been
activated and ELT is transmitting.
(4) Reset the G-switch, and restore other
disturbed switches to normal.
(5) Reinstall ELT in airplane.
e. Check calendar date for replacement of battery
pack. This date is supplied on a sticker attached to the
outside of the ELT case and to each battery.
17-127. REMOVAL AND NSTALLATIONOF
TRANSMITTER. (Refer to figure 17-22.)
a. Remove baggage curtain to gain access to thefigure
transmitter and antenna.
b. Disconnect coaxial cable from end of transmitter.
c. Cut sta-strap securing antenna cable and unlatch
metal strap to remove transmitter.
NOTE
Transmitter is also attached to the m
mounting bracket velcro strips; pull
transmitter to free from mounting bracket
and velcro.
NOTE
To replace velcro strips, clean surface
thoroughly with clean cloth saturated in
one of the following solvents: Trichloric
thylene, Aliphatic 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,
and 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
ensure intimate contact. Allow 24 hours
for complete cure.
17-68
Revision 3
d.
To reinstall transmitter, reverse preceding steps.
NOTE
An installation tool is required to properly
secure sta-strap. This tool may be
purchased locally or ordered from the
Panduit Corporation, Tinley Park, III,
Part No. GS-2B (conforms to MS90387-1).
CAUTION
Ensure that the direction of flight arrows
(placarded on the transmitter) are
pointing towards the nose of the aircraft.
17-128. REMOVAL AND INSTALLATION OF
ANTENNA. (Refer to figure 17-22.)
a. Disconnect coaxial cable (9) from base of antenna
(12).
Remove nut and lockwasher attaching antenna
base to fuselage, and the antenna (12) 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.
17-129. REMOVAL AND INSTALLATION OF
LITHIUM FOUR-CELL BATTERYPACK. (Referto
17-23
NOTE
Transmitters equipped with the 4-cell
battery pack can only be replaced with
another 4-cell battery pack..
NOTE
When existing battery fails or exceeds
normal expiration date, convert ELT
System to new D/M alkaline powered
ELT per Avionics Service Letter AV7831, dated November 20, 1978.
a. After the transmitter has been removed from
aircraft in accordance with paragraph 17-127, 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
gasket
and screws for
reinstallation.
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 17-23.
e. Connect the electrical connector as shown in figure
17-23
MODEL 210 &T210 SERIES SERVICE MANUAL
NOTE
NOTE
Before installing a new 4-cell battery
pack, check to ensure that its voltage is
11.2 volts or greater.
CAUTION
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.
If it is desirable 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.
f. Replace the transmitter cover and gasket.
g. Remove the old battery pack placard from end of
transmitter and replace with battery pack placard
supplied with the new battery pack.
BATTERY PACK
C589510-0210
TRANSMITTER
C589510-0209
WARNING
Figure 17-23. Lithium 4-Cell
The battery pack is pressurized contents.
Do NOT recharge, short circuit, dispose of
in fire of compact.
17-130. 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
Revision 3
17-68A/07.68B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
17-130. TROUBLE SHOOTING (Cont.)
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.
TROUBLE
*POWER LOW
REMEDY
PROBABLE CAUSE
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 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 750jackplug
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 current drain
is in the 85-100 ma range, the transmitter or
the coaxial cable is faulty.
Faulty coaxial
antenna cable.
4. Check coaxial antenna cable for high
resistancejoints. If this is found to be the case,
the cable should be replaced.
'This test should be carried out with the coaxial cable provided with your unit.
17-131. EMERGENCY LOCATOR TRANSMITTER.
BEGINNING WITH 21061716.
17-132. DESCRIPTION. The ELT is a self-contained,
solid state unit, having its own power supply with an
externally mounted antenna. 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 reception by
search and rescue personnel and others monitoring
the emergency frequencies. The ELT exhibits line
of sight transmission characteristics which correspond approximately to 100 miles at a search altitude
of 10,000 feet. The C589511-0103 transmitter, and
the C589511-0104 transmitter on aircraft with Canadian registry, are used thru 21062954. The C5895110117 transmitter, and the C589511-0113 transmitter
on aircraft with Canadian registry, are used on 21062955 thru 21064780. Beginning with 21064781 the
C589512-0103 transmitter is used on all aircraft.
The C589511-0104 transmits on 121. 5 MHz at 25 mw
rated power output for 100 continuous hours in the
temperature range of -40' to +131*F (-40-C to +55-C).
The C589511-0113 transmits on 121.5 MHz at 25 mw
rated power output for 100 continuous hours in the
temperature range of -4°F to +131*F (-20-C to +55ºC).
The C589511-0103 transmits on 121.5 and 243.0 MHz
simultaneously at 75 mw rated power output for 48
continuous hours in the temperature range of -40*F
to +131*F (-40ºC to +55-C). The C589511-0117 and
C589512-0103 transmits on 121. 5 and 243.0 MHz at
75 mw rated power output for 48 continuous hours in
the temperature range of -4°F to +131*F (-20*C to
+55-C).
Power is supplied to the transmitter by a battery pack.
The C589511-104 and C589511-0103 ELTs equipped
with a lithium battery pack must be modified by
SK185-20 as outlined in Avioncis Service Letter
AF78-31, dated 20 November 1981 to incorporate
Revision 3
17-69
I
MODEL 210 & T210 SERIES SERVICE MANUAL
alkaline battery packs. The C589511-0114 alkaline
battery packs have the service life of the battery pack
stamped on the battery pack, on the end of the
transmitter below the switch and on top of the
| transmitter. The C589512-0107 alkaline battery packs
have the replacement date and date of installation on
the top of the transmitter.
17-133. 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.
CAUTION
Do not leave the emergency locator
transmitter in the ON position longer
than 1 second (3 sweeps of the warble
tone) or you may activate downed
aircraft procedures by C.A.P., D.O.T., or
F.A.A. personnel.
17-134. OPERATIONAL TEST OF EMERGENCY
LOCATOR SYSTEM. The ELT, its battery pack, and its
antenna must be inspected and tested each 100 hours.
The operational test of the airplane's emergency locator
system should check both radiated signal strength and
the ELT G-switch. The airplane's VHF receiver is
located very close to the ELT and is very sensitive.
Consequently, using the airplane's VHF receiver to
monitor ELT transmission does not provide same level of
confidence in verifying ELT signal as using AM radio or
performing control tower check.
CAUTION
Tests with the antenna connected should
be approved by the nearest control tower.
The FAA/DOT allows free space
transmission tests from the airplane only
within first five minutes after each hour.
The test time allowed is limited to three
sweeps of the warble tone or
approximately one second. The control
tower should be notified that a test is
about to be conducted.
NOTE
After accumulated test or operation time
equals one hour, battery pack
replacement is required.
| a. Operational test of radiated signal with control
tower monitoring.
(1) Turn airplane master switch ON.
(2) Verify that test is conducted within first five
minutes of the hour.
(3) Turn airplane transceiver ON, request
permission from nearest control tower and flight service
17-70
Revision 3
station to conduct operational test of ELT, and request
control tower monitoring.
(4) Place ELT function selector to the ON position
for one second or less (no more than three sweeps of the
audio signal). Immediately replace the ELT function
selector to the ARM position after testing ELT.
(5) Contact control tower and confirm proper
locator beacon operation.
(6) Restore switches to normal.
b. Operational test of radiated signal with handheld
AM radio monitoring.
(1) Turn airplane master switch ON.
(2) Verify that test is conducted within first five
minutes of the hour.
(3) Turn airplane transceiver ON and request
permission from nearest control tower and flight service
station to conduct operational test of ELT.
(4) Position a small hand held AM radio tuned to
any frequency within six inches of the ELT antenna.
(5) Place ELT function selector to the ON position
for one second or less (no more than three sweeps of the
audio signal). Immediately replace the ELT function
selector to the ARM position after testing ELT.
(6) Verify that ELT signal has been detected on
hand held AM radio.
(7) Restore switches to normal.
c. Operational test of the TSO-C91 ELT G-switch.
(1) Remove ELT from airplane.
(2) While holding ELT in one hand, sharply strike
the end of the case in the direction of activation
indicated on the case of the transmitter.
(3) Using either radiated signal test method
described above, verify that the G-switch has been
activated and ELT is transmitting.
(4) Reset the G-switch, and restore other
disturbed switches to normal.
(5) Reinstall ELT in airplane.
d. Operational test of the TSO-C91a ELT G-switch.
(1) Remove ELT in airplane.
(2) While holding ELT firmly in one hand, make a
throwing motion followed by a sudden reversal of the
transmitter.
(3) Using either radiated signal test method
described above, verify that the G-switch has been
activated and ELT is transmitting.
(4) Reset the G-switch, and restore other disturbed
switches to normal.
(5) Reinstall ELT in airplane.
e. Check calendar date for replacement of battery
pack. This date is supplied on a sticker attached to the
outside of the ELT case and to each battery.
17-135. REMOVAL AND INSTALLATION OF
TRANSMITTER. (Refer to figure 17-24).
a. Remove baggage curtain to gain access to the
transmitter and antenna.
b. Disconnect coaxial cable from end of transmitter.
c. Remove the two #10 screws from the baseplate of
the ELT and remove ELT.
d. To reinstall transmitter, reverse preceding steps.
CAUTION
Ensure that the direction of flight arrows
(placarded on the transmitter) are
pointing towards the nose of the aircraft.
MODEL 210 & T210 SERIES SERVICE MANUAL
II.1
CIItLt0*
PLACARD LOCATED ON UPPER R.H. ^::-"'
.""-~'~-
10
I
-i~
2
I
- ........
/ ; ; ^ >^.....
.' .....- ...... . ... .. . .
:
APPLIES TO AIRCRAFT
WITH PITCH ACTUATOR
BEGINNING WITH
21062828
1
ELT IS LOCATED BEHIND
THIS SURFACE
3
PLACARD LOCATED ON RIGHT
HAND SIDE OF TAILCONE ADJACENT
TO ELT. ON CANADIAN AIRCRAFT.
3
Detail A
BEGINNING WITH 21064781
Detail C
ROTATED 180 °
Figure 17-24.
Emergency Locator Transmitter Installation (Sheet 3 of 3)
17-73
MODEL 210 & T210 SERIES SERVICE MANUAL
17-136. REMOVAL ANDINSTALLATIONOF
ANTENNA (Refer to figure 17-24.)
a. Disconnect coaxial cable from base of antenna.
b. Remove the nut and lockwasher attaching the
antenna base to the fuselage and the antenna will be
free for removal.WARNING
c. To reinstall the antenna, reverse the preceding
steps.
CAUTION
The C589511-0111 and C589511-0119 coaxial cable must be installed as indicated
on the cable sleeve. Cable end marked
"TO ANT" must be connected to the ELT
antenna, and the end marked "TO ELT"
must be connected to the C589511-0113/
-0117 and C589511-0103/-0104 transmitters.
g. Stamp the new replacement date on the outside
of the ELT. The date should be noted on the switching nameplate on the side of the unit as well as on the
instruction nameplate on top of the unit.
WARNINI
The battery pack has pressurized
WARNINGcontents.
Do not recharge, short circuit, or dispose of in
fire.
CAUTION
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.
DO NOT use a substitute battery pack.
NOTE
Upon reinstallation of antenna, cement
rubber boot (14) using RTV 102, General
Electric Co. or equivalent, to antenna
whip only; do not apply adhesive to fuselage skin or damage to paint may result.
C589511-0103 TRANSMITTER
C589511-0104 TRANSMITTER (CANADIAN)
17-137. REMOVAL AND INSTALLATION OF
BATTERY PACK (See figure 17-25).
NOTE
Transmitters equipped with
C589511-0105 or C589511-0106 battery
packs can be replaced with a
C589511-0114 after modification by
SK185-20 has been completed.
CAUTION
Lithium battery pack must be replaced
with alkaline battery packs per SK185-20.
a. After the transmitter has been removed from
aircraft in accordance with paragraph 17-135, place the
transmitter switch in the OFF position.
b. Remove the four screws attaching the cover to the
case and then remove the cover to gain access to the
battery pack.
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 17-25.
e. Connect the electrical connector as shown in figure
17-25.
C589511-0105 BATTERY PACK
C589511-0106 BATTERY PACK
(CANADIAN)
C589511-0117 TRANSMITTER
C589511-0113 TRANSMITTER (CANADIAN)
Before installing the C589511-0105 pack,
check to ensure that its voltage is 7. 5
volts or greater.
f. Replace the transmitter baseplate on the unit and
pressing the baseplate and unit together attach baseplate with four nylok patch screws.
C589511-0114 DOMESTIC &
CANADIAN
Figure 17-25. Battery Pack Installation.
17-74
Revision 3
MODEL 210 & T210 SERIES SERVICE MANUAL
17-138. 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
TROUBLE
*POWER LOW
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.
REMEDY
PROBABLE CAUSE
Low battery voltage.
1. Settoggleswitchtooff.
2. Disconnect the battery pack from the
transmitter and connect a Simpson 260 model
voltmeter and measure voltage. If the battery
pack transmitters is 7.5 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. Reconnect battery pack to the
transmitter.
b. Using E. F. Johnson 105-0303-001
jackplugs and 3-inch maximum long leads,
connect a Simpson Model 1223 ammeter to the
jack.
c. Set the toggle switch to AUTO and
observe the ammeter current drain. If the
current drain is in the 15-25 ma range, the
transmitter or the coaxial cable is faulty.
Faulty coaxial
antenna cable.
4. Check coaxial 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 coaxial cable provided with your unit.
Revision 3
17-75
MODEL 210 & T210 SERIES SERVICE MANUAL
OPTIONAL EQUIPMENT (RUNNING LOAD)
Cessna 400B Nav-O-Matic (Type AF-550A)
(Includes Unslaved DG GS-502A)
With Slaved Directional Gyro System
.
.
1977
1978
1979
5. 0
5.0
5.0
5. 2
5. 2
5. 4
5. 4
5. 8
6. 0
6. 0
AMPS
1980
1981
1982
1983
5. 0
5. 0
5.0
5.0
5. 2
5.2
5. 2
5.2
5.2
5.4
5. 4
5. 4
5. 4
5.4
5.5
5.5
5.5
5.3
5. 3
5.3
5.8
5.8
.
5.8
5.8
5. 8
6. 0
6. 0
6.0
6. 0
6. 0
1.0
1.0
1.0
1.0
0. 5
1.0
1.50
1.00
1.0
0. 5
1.2
0.65
0.1
0.1
0.10
0.10
(CS-504A)
With Slaved D.G. & Course Datum . ...
(CS-504A)
With Unslaved HSI (IG-832C) .
.......
With Slaved HSI System (CS-832A) . . . .
With Slaved HSI & Course Datum .....
(CS-832A)
Stereo Avionics West ..........
Cessna 400 DME (RT-477A) .......
Cessna 400 R-Nav (RN-479A) . ....
EC-100 Stereo ...............
Bonzer Radar Altimeter ..........
DME - 190
.
.................
2.9
DME-451 ................
ANS - 351 RNAV ..............
Altitude Encoder (Blind) ..
..
.0.1
High Altitude Encoder (Blind) ........
De-Ice System (Certified for Flight in
Icing Conditions) .............
Prop De-Ice ................
Windshield De-Ice .............
RDR-160 Weather Radar ..........
Cessna 400 Glideslope (Type R-443B) .....
Cessna 400 Nav/Com (Type RT-428A). ....
HSI System (IG-832A) ...........
Slaved Directional Gyro (G-504A).
1.2
0.65
...........
0.1
3. 5
9. 5
0. 32
1. 5
8 or *1.
46.2
46.2
40.65
3. 5
0. 5
3.5
0. 5
3.5
0. 5
3.5
0. 5
18.0
16.0
3.5
0. 5
1.0
0.35
1.0
.35
1.0
0.4
1.0
0.4
1.0
0.4
1.0
2. 3
2.3
2.3
2. 30
2. 3
4.
7.5
9.
3.0
7.0
10.0
3.
.28
4.
7. 5
9.
3. 0
7.0
10.0
3.6ea
.28
2. 3
7.
2.
7.5
2.3
7.5
3.0
7.0
8. 5
3. 6ea
.28
3. 0
3. 0
8. 5
3. 6ea
1.8
3. 6ea
.40
. . . . . . 15 or *. 3
Foster RNAV 511
.............
400 RMI
.
.
...............
Avionics Cooling Fan ............
Interphone System ..............
ITEMS NOT CONSIDERED AS PART OF
RUNNING LOAD
Cessna 300 Nav/Com (RT-385A)......
Cessna 400 Nav/Com (RT-485A, RT-485B) . .
ASB-125 SSB HF Transceiver ....
.
PT-10A Transceiver .......
.....
Auxiliary Fuel Pump ..
. . . . ....
Cigarette Lighter ..............
Flap Motor ................
Landing Lights (Each) ............
Stall Warning Horn .............
Wing Courtesy Lights and Cabin Lights . ...
Ice Detector Light .............
Hydraulic Power Pack ...........
2.
7. 5
3. 0
7.0
8.5
.28
30
4.0
7.5
0 9.
3.0
7.0
10.0
3.
3.57
6e
.28
1.2
1.2
1.2
1.2
1.43
1.5
1.5
1.5
8.0
8.00 17.5
17.5
40.00 40.00
0.
.7
0. 7
0. 7
0.1
1 0.
1 0. 0.1
5. 35
Electric Elevator Trim ...........
.
Map Light (Glare Shield or Control Wheel). .
Recognition Lights ............
Air Conditioning ..............
*Console Lights not used with post lights.
*Only one or the other may be used at one time.
tNegligible
*In flight running load
*With Bootstrap OTransmit
*Receive
1.2
1. 5
1.5
17.5
1.5
17.5
1.5
14.0
0. 7
0. 1
5.3
22.8
0. 7
0.1
5.3
22.8
.51
0.1
3. 6
22.8
Minimum to Maximum
17-77
MODEL 210 & T210 SERIES SERVICE MANUAL
ELECTRICAL LOAD ANALYSIS CHART
AMPS
1984
STANDARD EQUIPMENT (RUNNING LOAD)
Battery Contactor .................................
Clock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cylinder Head Temperature ................
.........
.................
Fuel Quantity Indicators .
........
Engine Instruments
......
.........
.................
...
.........
Flashing Beacon .......
Instrument Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..
a. Electroluminescent Panel..............................
b. Cluster . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
c. Console * .
. . ..
. . ..
. . . . ... . . . . . ...... . . ...
Instrument Lights .....
...............
.2.2
EL Panels
....................
..................
............
.............
Solenoid Valve - Gear Doors & Warning.
Lamp - Gear Up or Gear Down .
..................
...
.....
Solenoid Valve - Gear Handle Lock ............................
Position Lights .....................................
Turn Coordinator ....................................
Turn & Bank Indicator (Optional) .....................
....
Alternator Control Unit ........
. . . .....
.....
..
0. 33
. . . . .
. .
.05
.10
.
.
.7.0
. . . .
. . . . .
. . . .
0.3
.
.
..
0.04
...
. ....
2.8
0.30
0.24
2.0
OPTIONAL EQUIPMENT (RUNNING LOAD)
Heated Pitot and Stall Warning Heaters ...................
....
. . ..
. ..
. . . . . . . . ..
. ..
..
. . . . ..
Windshield Anti-Ice
Wing De-Ice ...........
........
.......................
Propeller Anti-Ice ..........
...................
..
.. .............
Strobe Lights ..
....
Post Lights ..
. . .......................
Cessna 200A Navomatic (Type AF-295B) ..
. . ..................
Cessna 300 ADF (Type R-546E) .............................
Cessna 300 Nav/Com (360 Channel-Type RT-308C)............
Cessna 300 Nav/Com (Type RT-328T) .............
Cessna 300 Transponder (Type RT-359A) ........................
Cessna 300A Navomatic (Type AF-395A) ........................
With Unslaved HSI (IG-832C).........................
Cessna 300 Nav/Com (RT-385A) ........................
Cessna 400 R-Nav (RN-478A) ..............................
Cessna 400 ADF (Type R-446A) ................
Cessna 400 Nav/Com (RT-485A, RT-485B) ..
. . .................
.........
Cessna 400 Transponder (RT-459A). ...................
Cessna 400 DME (Type RT-476A, Type 478A) .......................
.......................
Cessna 400 Encoding Altimeter (EA-401A) .
Bendix GM-247A Marker Beacon .............................
Cessna 400 Marker Beacon (Type R-402A, ........................
R-402 B)
Sunair SSB Transceiver (Type ASB-125). .........................
Pantronics PT-10A HF Transceiver ...........................
Altitude Alert/Select (AA-801A) ............................
Cessna 400 Nav-O-Matic (Type AF-420A) ...............
..........
With Slaved Directional Gyro System .........................
Cessna 400B IFCS (Type IF-550A) (Includes
HSI & Course Datum) .................
................
17-78
Revision 2
...
5. 8
4.4
3.0
18.0 ·
2.0
0. 8
2.5
1.0
.
....
.
.
2.0
2.5
2.
1.0*
1.6
1. 6*
2.0
0.1
0.1
2.5 *
.
0.
6.0
B
MODEL 210 & T210 SERIES SERVICE MANUAL
AMPS
1984
OPTIONAL EQUIPMENT (RUNNING LOAD)
Cessna 400B Nav-O-Matic (Type AF-550A) ...................
. ....
(Includes Unslaved DG GS-502A)
With Slaved Directional Gyro System
.........................
(CS-504A)
With Slaved D. G. & Course Datum ..........................
(CS-504A)
WithUnslaved HSI (IG-832C).
. ...
.....
..
.
.
..............
With Slaved HSI System (CS-832A) .....................
With Slaved HSI & Course Datum ...........................
(CS-832A)
Cessna 400 DME (RT-477A) ..
.
..........................
Cessna 400 R-Nav (RN-479A) .....................
.......
EC-100 Stereo ... . ..
. . . . . . ...
. . . . . . . . . . . . . ..
. ..
...
Bonzer Radar Altimeter..
....
...
...........
DME - 190 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
DME - 451 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.0
5.2
5. 3
5.8
6.0
1. 50
1.00
1.0
0. 5.
ANS- 351 RNAV ....................................
Altitude Encoder (Blind) ......................
High Altitude Encoder (Blind) ................
De-Ice System (Certified for Flight in
Icing Conditions) . ..
. . . . . . . . . . . . . . . . . . . . . . . . .
Prop De-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Windshield De-Ice ..............
..
.......
RDR-160 Weather Radar ................................
Cessna 400 Glideslope (Type R-443B) ..........................
Cessna 400 Nav/Com (Type RT-428A) ..
. . ...............
HSI System (IG-832A) .. . . . . . . . . . ..
. . . . . . ..
. . . ...
Slaved Directional Gyro (G-504A) ..............
Foster RNAV 511 ..................................
400 RMI
... ........................
...........
Avionics Cooling Fan ..................................
Interphone System ...................................
Primus 100 WX Radar . .. . .............................
King KRA-10A Radar Altimeter ...............
..
............
King KMA-24-03 Audio Panel W/MKR ....................
.
King KX-165 Nav/Comm W/GS. ..
. . . ...........
King KY-196 Comm Transceiver .............................
King KNS-81 RNAV/G. S.
....
. ...
...................
King KN-63 DME ..........
...................
King K12-87 ADF..........
...............................
King KT-79 Transponder ..
.
....................
King KI-229 RMI .....
............
..
..........
King KT-98 Radio Telephone ......................
.
King KWX-56 Color WX Radar ...................
..
0. 10
0.10
. . . . . . .
. .
......
40. 65
.
..
3.5
0.5
. . . . .
0. 4
1.0
†
.
...
. ...
2.0
0. 20
0.16
0.4
0.4
.50
.60
.43
.36
1.00
0.....50
3.0
.......
.....
.........
ITEMS NOT CONSIDERED AS PART OF
RUNNING LOAD
Cessna 300 Nav/Com (RT-385A) .............
Cessna 400 Nav/Com (RT-485A, RT-485B)
ASB-125 SSB HF Transceiver .......................
PT-10A Transceiver . . . . . . . . . .
Auxiliary Fuel Pump .
.
..
Cigarette Lighter . . . . . . . . . . .
Flap Motor .
............................
Landing Lights (Each) . .. . . . .......
Stall W arning Horn . . . . . . . . . . .
.
2.3
2. 3
7.5
....
. . . . . . . . . . . . . . . . . . . . . . . .
.................
....
....
. . . . . . . . . . . . . . . . . . . . . . . .
. . .....
. . .....
..
. .......
. ..
. . . . . . . . . . . . . . . . . . . . . . . .
Revision 2
3.0
1.8
3.6 EA
40
17-79
MODEL 210 & T210 SERIES SERVICE MANUAL
AMPS
1984
ITEMS NOT CONSIDERED AS PART OF
RUNNING LOAD
Wing Courtesy Lights and Cabin Lights ............
Ice Detector Light .........
.......................
Hydraulic Power Pack .
............................
..............
..
...
Electric Elevator Trim .......
......................
.
Map Light (Glare Shield or Control Wheel) ........................
Recognition Lights ........
.....
. . ........
.......
..
Air Conditioning .......
...................
King KX165 . . . . . . . . . .
.. . . . . . . . . . . .
King KY196.
.............
.
.
.
.
.
KT-96 Radio Telephone ...
. . . . . . .
..................
*Console Lights not used with past lights.
*Only one or the other may be used at one time.
tNegligible
*In night running load
*With Bootstrap O Transmit
17-80
Revision 2
1.5
1.5
14.0
.
*Receive
...
.
.
..... .
51
0.1
3. 6
22.8
4.5
5.0
3.0
0 Minimum to Maximum
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 18
STRUCTURAL REPAIR
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
3D12/18-2
........
STRUCTURAL REPAIR
3D12/18-2
Repair Criteria .........
3D12/18-2
Equipment and Tools .......
......
. 3D12/18-2
Support Stands
. .3D12/18-2
Fuselage Repair Jigs .
.. 3D12/18-2
Wing Jigs .......
. 3D12/18-2
Repair Materials ......
Wing Twist and Stabilizer
. .. 3D12/18-2
Angle-of-Incidence ....
3D12/18-2
Wing .............
. 3D12/18-2
Description ......
3D12/18-2
.......
Wing Skin .
. 3D12/18-2
Negligible Damage ...
Repairable Damage . . . . 3D12/18-2
Damage Necessitating Replace3D13/18-3
ment of Parts ......
3D13/18-3
Wing Stringers ........
3313/18-3
Negligible Damage .....
.....
3D13/18-3
Repairable Damage
Damage Necessitating Replace3D13/18-3
ment of Parts ......
3D13/18-3
Wing Ribs ..........
.3D13/18-3
Negligible Damage ...
. 3D13/18-3
Repairable Damage ...
Damage Necessitating Replace. 3D13/18-3
ment of Parts ....
.3D13/18-3
........
Wing Spar
3D13/18-3
Negligible Damage .....
. 3D13/18-3
Repairable Damage .
Damage Necessitating Replace.3D13/18-3
ment of Parts .....
Wing Fuel Bay Spars and Ribs . 3D13/18-3
.3D13/18-3
Negligible Damage ...
. 3D13/18-3
Repairable Damage ...
Damage Necessitating Replace.3D13/18-3
ment of Parts .....
........
. 3D13/18-3
Ailerons .
3D13/18-3
Negligible Damage .....
. 3D13/18-3
Repairable Damage ...
Damage Necessitating Replace. 3D13/18-3
ment of Parts .....
3D13/18-3
Aileron Balancing .....
. 3D14/18-4
Wing Flaps .........
3D14/18-4
Negligible Damage .....
3D14/18-4
Repairable Damage .....
Damage Necessitating Replace3D14/18-4
ment of Parts .....
3D14/18-4
Wing Leading Edge ......
3D14/18-4
Negligible Damage .....
3D14/18-4
Repairable Damage .....
Damage Necessitating Replace.3D14/18-4
.....
ment of Parts
.....
3D14/18-4
Elevators and Rudder
. 3D14/18-4
Negligible Damage ....
. 3D14/18-4
Repairable Damage ....
Damage Necessitating Replace3D14/18-4
.....
ment of Parts
Elevator and Rudder
.3D14/18-4
.......
Balancing
3D14/18-4
Fin and Stabilizer ......
.3D14/18-4
Negligible Damage ....
....
3D14/18-4
Repairable Damage
Damage Necessitating Replace...
.3D14/18-4
ment of Parts
.
.3D14/18-4
.
....
Bonded Doors
3D14/18-4
Repairable Damage ...
3D14/18-4
..
......
Fuselage
.3D14/18-4
Description .......
3D14/18-4
Negligible Damage .....
. 3D15/18-5
Repairable Damage ....
Damage Necessitating Replace3D15/18-5
......
ment of Parts
3D15/18-5
Bulkheads ..........
Landing Gear Bulkheads . . 3D15/18-5
Repair After Hard Landing. 3D15/18-5
. 3D15/18-5
Firewall Damage .....
3D15/18-5
..........
Fasteners
.3D15/18-5
Rivets ..........
I
Replacement of Hi-Shear
. 3D15/18-5
Rivets. .........
3D16/18-6A
Substitution of Rivets ...
3D19/18-6D
Baffles ...........
3D19/18-6D
Engine Cowling .......
Repair of Cowling Skins . . 3D19/18-6D
Repair of Reinforcement
.3D20/18-7
Angles .......
Repair of Glass-Fiber
Constructed Components . . 3D20/18-7
Revision 3
18-1
MODEL 210 & T210 SERIES SERVICE MANUAL
18-1.
IR
REPA
STRUCTURAL
18-2. REPAIR CRITERIA. 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. 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 suggests the extent of structural repair practicable on the aircraft and supplements Federal Aviation Regulation, Part 43. Consult the factory when in doubt about a repair not
specifically mentioned here.
18-3. EQUIPMENT AND TOOLS.
18-4. SUPPORT STANDS. Padded, reinforced sawhorse or tripod type support stands, sturdy enough to
support any assembly placed upon them, must be used
to store a removed wing or tailcone. Plans for local
fabrication of support stands are contained in figure
18-1. The fuselage assembly, from the tailcone to the
firewall, must NOT be supported from the underside,
since the skin bulkheads are not designed for this purpose. Adapt support stands to fasten to the wingattach points or landing gear attach-points when supporting a fuselage.
18-5. FUSELAGE REPAIR JIGS. Whenever a repair
is to be made which could affect structural alignment,
suitable jigs must be used to assure correct alignment of major attach points, such as fuselage, firewall, wing and landing gear. These fuselage repair
jigs are obtainable from the factory.
18-6. WING JIGS. These jigs serve as a holding
fixture during extensive repair of a damaged wing,
and locates the root rib, leading edge and tip rib of
the wing. These jigs are also obtainable from the
factory.
18-7. REPAIR MATERIALS. Thickness of a 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 readily be determined, 2024T3 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 gauge of
the material being repaired unless otherwise noted.
18-2
It is often practical to cut repair pieces from service
parts listed in the Parts Catalog. A few components
(empennage tips, for example) are fabricated from
thermo-formed plastic or glass fiber constructed
material.
18-8. WING TWIST AND STABILIZER ANGLE-OFINCIDENCE. Wing twist (washout) and stabilizer
angle of incidence are shown below. Stabilizers do not
have twist. The cantilever wing has a uniform twist
from the root rib to the tip rib. Refer to figure
18-2 for wing twist measurement.
18-9.
WING
Twist (Washout)
3°
STABILIZER
Angle-of-incidence
-3°± 15'
WING.
18-10. DESCRIPTION. The wing is sheet-metal
constructed, with a single main spar, two fuel spars,
formed ribs and stringers. The front fuel spar also
serves as an auxiliary spar and is the forward wing
attaching point. An inboard section forward of the
main spar is sealed to form an integral fuel bay
area. The main spar consists of milled spar caps
and attaching fittings joined by a web section. The
aft fuel spar is a formed channel. The front fuel
spar is a built-up assembly consisting of a formed
channel, doubler, attach strap and support angle.
Stressed skin, riveted to the ribs, spars and stringers, completes the wing structure. Access openings
(hand holes with removable cover plates) are located
in the underside of the wing between the wing root
and tip section. These openings afford access to the
flap and aileron bellcranks, flap drive pulleys, flap
actuator in left wing, flap and aileron control cable
disconnect points, fuel adapter plate, air scoop connectors and electrical wiring.
18-11.
WING SKIN.
18-12. 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 any area of the wing, Outboard of wing station
40.00 in areas of low stress intensity, cracks, deep
scratches or 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.
The area on the lower surface of the wing between
the two stringers adjacent to the main spar is not
considered low stress intensity. Stop drilling is
considered a temporary repair and a permanent repair should be made as soon as practicable.
18-13. REPAIRABLE DAMAGE. Repairs must not
be made to the upper or lower wing skin inboard of
station 40. 00 without factory approval. However, an
.
MODEL 210 & T210 SERIES SERVICE MANUAL
entire skin may be replaced without factory approval.
Refer to Section 1 for wing station locations. Figure
18-4 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 one-half 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
fush 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 opoxy type filler may be used at such
joints.
scratches and abrasions may be considered negligible.
18-25. REPAIRABLE DAMAGE. All cracks, stress
wrinkles, deep scratches and sharp dents must be
repaired. However, repairs must not be made to
the main wing spar inboard of wing station 155.00
without factory approval. Refer to Section 1 for wing
station locations. Figure 18-7 outlines a typical
main wing spar repair.
18-26. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. An entire wing spar may be replaced
without factory approval.
18-27. WING FUEL BAY SPARS AND RIBS.
18-14. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If a skin is badly damaged, repair must
be made by replacing an entire skin panel, from one
structural member to the next. Repair seams must
be made to lie along existing structural members
and each seam must 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
must be copied. If the repair ends at a structural
member where no seam is used, enough repair panel
must be used to allow an extra row of staggered
rivets, with sufficient edge margin, to be installed.
18-28. NEGLIGIBLE DAMAGE. Any smooth dents
in the fuel spars that are free from cracks, abrasions and sharp corners, which are not stress
wrinkles and do not interfere with any internal struc ture or mechanism, may be considered as negligible
damage in any area of the spar.
18-15. WING STRINGERS.
18-30. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Due to the amount of fuel bay sealant
which must be removed from fuel bay components to..
facilitate repair, individual parts are not available
to replace fuel bay spars or ribs. The entire fuel
bay area must be replaced as a unit.
18-16. NEGLIGIBLE DAMAGE.
18-12.
Refer to paragraph
18-17. REPAIRABLE DAMAGE. Figure 18-5 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.
18-18. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If a stringer is so badly damaged that
more than one section must be spliced, replacement
is recommended.
18-19. WING RIBS.
18-20. NEGLIGIBLE DAMAGE. Refer to paragraph
18-12.
18-21. REPAIRABLE DAMAGE.
trates typical wing rib repairs.
Figure 18-6 illus-
18-22. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Any wing rib damaged extensively
should be replaced. However, due to the necessity
of disassembling so much of the wing in order to replace a rib, especially in the fuel bay area which involves sealing, wing ribs should be repaired if practicable.
18-23.
WING SPAR.
18-24. NEGLIGIBLE DAMAGE.
Due to the stresses
which the wing spar encounters, very little damage
can be considered negligible.
Smooth dents, light
18-29. REPAIRABLE DAMAGE. The type of repair
outlined in figure 18-7 also applies to fuel bay spars
outboard of wing station 124.0. Inboard of station
124. 0, factory approval of proposed repairs is required. Refer to Section 13 for sealing procedures
when working in fuel bay areas.
18-31. AILERONS.
18-32. NEGLIGIBLE DAMAGE.
18-12.
Refer to paragraph
18-33. REPAIRABLE DAMAGE. The repair shown
in figure 18-8 may be used to repair damage to aileron leading edge skins. The flush-type skin patches
shown in figure 18-4 may be used to repair damage
to the remaining skins. Following repair, the aileron must be balanced. Refer to paragraph 18-35 and
figure 18-3 for balancing the aileron.
18-34. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If the damage would require a repair
which could not be made between adjacent ribs, complete skin panels must be replaced. Ribs and spars
may be repaired, but replacement is generally pre ferable. Where extensive damage has occurred, replacement of the aileron assembly is recommended.
After repair or replacement, balance aileron in
accordance with paragraph 18-35 and figure 18-3.
18-35. AILERON BALANCING. Following repair,
replacement or painting, the aileron must be balanced.
A flight control surface balancing fixture kit is avail-
able (P/N 5180002-1).
See figure 18-3 for procedures
pertaining to the use of this kit.
Revision 2
18-3
MODEL 210 & T210 SERIES SERVICE MANUAL
18-36.
WING FLAPS.
18-37.
18-12.
NEGLIGIBLE DAMAGE.
Refer to paragraph
18-38. REPAIRABLE DAMAGE. Flap repairs should
be similar to aileron repairs discussed in paragraph
18-33. A flap leading edge repair is shown in figure
18-9.
18-39. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Flap repairs which require replacement
of parts should be similar to aileron repairs discussed in paragraph 18-34. Since the flap is not considered a movable control surface, no balancing is
required.
18-40.
WING LEADING EDGE.
18-41.
18-12.
NEGLIGIBLE DAMAGE.
Refer to paragraph
18-42. REPAIRABLE DAMAGE. A typical leading
edge skin repair is shown in figure 18-8. Also, wing
skin repairs, outlined in paragraph 18-13, may be
used to repair leading edge skins, although the flushtype patches should be used. Extra access holes,
described in figure 18-10, must not be installed in
the wing without factory approval. Where extreme
damage has occured, replace complete skin panels.
18-43. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. An entire leading edge skin may be replaced without factory approval.
18-48. ELEVATOR AND RUDDER BALANCING.
Following repair, replacement or painting, the
elevators and rudder must be balanced. A flight
control surface balancing fixture kit is available
(P/N 5180002-1). See figure 18-3 for procedures
pertaining to the use of this kit.
18-49.
18-50.
18-12.
18-52. 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 must be replaced.
Ribs and spars may be repaired, but replacement is
generally preferable. Where damage is extensive,
replacement of the entire assembly is recommended.
BONDED DOORS.
ELEVATORS AND RUDDER.
18-45. NEGLIGIBLE DAMAGE. Refer to paragraph
18-12. The exception to negligible damage on the
elevator surfaces is the front spar, where a crack
appearing in the web at the hinge fittings or in the
structure 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.
18-52B. 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.
18-53.
18-46. REPAIRABLE DAMAGE. Skin patches
illustrated in figure 18-4 may be used to repair skin
damage. Following repair, the elevators and rudder
must be balanced. Refer to paragraph 18-48 and
figure 18-3 for balancing the elevators and rudder.
If damage would require a repair which could not be
made between adjacent ribs, see the following paragraph.
18-47. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. If the damaged area would require a repair which could not be made between adjacent ribs,
complete skin panels must be replaced. Ribs and
spars may be repaired, but replacement is generally
preferable. Where extensive damage has occured,
replacement of the entire assembly is recommended.
After repair and/or replacement, balance elevators
and rudder in accordance with paragraph 18-48 and
figure 18-3.
18-4
Refer to paragraph
18-51. REPAIRABLE DAMAGE. Skin patches illustrated in figure 18-4 may be used to repair skin
damage. Access to the dorsal area of the fin may be
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 orone 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 would be located in an area with
compound curves, see the following paragraph.
18-52A.
18-44.
FIN AND STABILIZER.
NEGLIGIBLE DAMAGE.
Revision 2
FUSELAGE.
CAUTION
Repairs must not be made to the main wing
spar carry-thru section of the cantilever
wing without factory approval.
18-54. DESCRIPTION. The fuselage is of semimonocoque construction consisting of formed bulkheads, longitudinal stringers, reinforcing channels
and skin platings.
18-55. NEGLIGIBLE DAMAGE. Refer to paragraph
18-12. Mild corrosion appearing upon alclad surfaces does not necessarily indicate incipient failure
of the base metal. However, corrosion of all types
must be carefully considered and approved remedial
action taken. Small cans appear in the skin structure
of all metal aircraft. It is strongly recommended,
MODEL 210 & T210 SERIES SERVICE MANUAL
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 occuring over
stringers which disappear when the rivet pattern is
removed may be considered negligible. However,
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 occuring in the skin of the main
landing gear bulkhead areas must not be
considered negligible. The skin panel must
be opened sufficiently to permit a thorough
examination of the lower portion of the landing gear bulkhead and its tie-in structure.
checked for alignment and a straightedge must be
used to determine deformation of the bulkhead webs.
Damaged support structure, buckled floorboards
and skins and damaged or questionable forgings
must be replaced.
18-61. FIREWALL DAMAGE. Firewalls may be repaired by removing the damaged material and splicing in a new section. The new portion must be lapped over the old material, sealed with Pro-Seal #700
(Coast Pro-Seal Co., Chemical Division, 2235 Beverly Blvd., Los Angeles, California) compound, or
equivalent and secured with MS16535 (steel) or MS20613 (corrosion-resistant steel) rivets. The heater
valve assembly is attached with MS16535 and MS20613 rivets. Firewall plates, firewall doublers,
and nutplates are attached to the firewall with MS20470 (aluminum) rivets. Damaged or deformed
angles and stiffeners may be repaired as shown in
figure 18-11, or they may be replaced. A severely
damaged firewall must be replaced as a unit.
Wrinkles occuring 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 must be identical to the existing manufactured seam at the edge of
the sheet.
18-62. FASTENERS. Fasteners used in the aircraft
are generally solid aluminum rivets, blind rivets, and
steel-threaded fasteners. Usage of each is primarily
a function of the loads to be carried, accessibility,
and frequency of removal. Rivets used in aircraft
construction are usually fabricated from aluminum
alloys. In special cases, monel, corrosion-resistant
steel and mild steel, copper, and iron rivets are used.
18-56. REPAIRABLE DAMAGE. Fuselage skin repairs may be accomplished in the same manner as
wing skin repairs outlined in paragraph 18-13.
Stringers, formed skin flanges, bulkhead channels
and similar parts may be repaired as shown in figure 18-5.
18-63. RIVETS. Standard solid-shank MS riets
are those generally used in aircraft construction.
They are fabricated in the following head types:
roundhead, flathead, countersunk head, and brazier
head. Flathead rivets are generally used in the aircraft interior where head clearance is required.
MS20426 countersunk head rivets are used on the
exterior surfaces of the aircraft to minimize turbulent airflow. MS20470 brazier head rivets are used
on the exterior surfaces of the aircraft where strength
requirements necessitate a stronger rivet head than
that of the countersunk head rivet. Both the brazier
head and the countersunk head rivets are used on the
exterior of the aircraft where head clearance is required. Hi-shear rivets are special, patented rivets
having a hi-shear strength equivalent to that of standardAN bolts. They are used in special cases in
locations where hi-shear loads are present, such as
in spars, wings, and in heavy bulkhead ribs. This
rivet consists of a cadmium-plated pin of alloy steel.
Some have a collar of aluminum alloy. Some of these
rivets can be reaily identified by the presence of the
attached collar in place of the formed head on standardrivets. Blind rivets are used, where strength
requirements permit, where one side of the structure
is inaccessible, making it impossible or impractical
to drive standard solid-shank rivets.
18-57. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Fuselage skin major repairs may be
accomplished in the same manner as wing skin repairs outlined in paragraph 18-14. Damaged fittings
must be replaced.
18-58.
BULKHEADS,
18-59. 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 must be repaired by replacing the landing gear support assembly as an aligned unit.
18-60. 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 must be carefully examined and
all support forgings must be checked for cracks,
using a dye penetrant and proper magnification.
Bulkheads in the area of possible damage must be
18-64. REPLACEMENT OF HI-SHEAR RIVETS.
Replacement of hi-shear rivets 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
18-5/(18-6 blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
using the following fasteners.
a. NAS464P-* bolt, MS21042-* nut and AN960-*
washer in place of Hi-shear rivets for forgings with
machined flat surfaces around attachment holes.
b. NAS464P-* bolt, ESNA2935-* mating base washer
for forgings
and ESNA RM52LH2935-* self-aligning nut
°
(with draft angle of up to a maximum of 8 ) without
machined flat surfaces around attachment holes.
*Dash numbers to be determined according to the size
of the holes and the grip lengths required. Bolt grip
length should be chosen so that no threads remain in
the bearing area.
18-65. SUBSTITUTIN OF RIVETS.
a. Solid-shank rivets (MS20426AD and MS20470AD).
Replace
In thickness
(or thicker)
When placing rivets in installations which require
raised head rivets, it is desirable to use rivets identical to the type of rivet removed. Countersunk-head
rivets (MS20426) are to be replaced by rivets of the
same type and degree of countersink. When rivet
holes become enlarged, deformed, or otherwise
damaged, use the next larger size rivet as a replacement. Replacement shall not be made with rivets of
lower strength material/
b. Hi-shear Rivets. When hi-shear rivets are not
available, replacement of sizes 3/16-inch or greater
rivets shall be made with bolts of equal or greater
strength than the rivet being replaced, and with selflocking nuts of the same diameter.
c. The following pages contain approved solid-shared
and hi-shear rivet substitutions.
With
MS20470AD3
.025
.020
NAS1398B4, NAS1398D4
NAS1738B4, NAS1738D4, NAS1768D4,
CR3213-4, CR3243-4
MS20470A04
.050
.040
NAS1398B4, NAS1398D4
NAS1398B5, NAS1398D5, NAS1738B4,
NAS1738E4, NAS1768D4, CR3213-4
NAS1738B5, NAS1738E5, NAS1768D5,
~CR3213-5, CR3243-4
CR3243-5
.032
.025t
.025
MS20470AD5
.063
.050
.040
.032
NAS1398B5, NAS1398D5
NAS1398B6, NAS1398D6, NAS1398B5,
NAS1738E5, CR3213-5
NAS1738B6, NAS1738E6, NAS1768D5,
CR3213-6, CR3243-5
CR3243-6
.050
NAS1398B6
NAS1398D6
NAS1738B6, NAS1738D6, NAS1768D6,
CR3213-6
CR3243-6
MS20426AD3
(Countersunk)
.063
.040
NAS1399B4, NAS1399D4
NAS1769D4, CR3212-4
(See Note 1)
.025
NAS1769B4, NAS1739E4, CR3242.4
MS20470AD6
.080
.071
.063
Revision 2
18-6A
MODEL 210 & T210 SERIES SERVICE MANUAL
Replace
With
In thickness
(or thicker)
MS20426AD4
(Countersunk)
.080
.063
.050
.040
NAS1399B4, NAS1399D4
NAS1739B4, NAS1739D4, CR32124
NAS1769D4
CR3242-4
(See Note 1)
.050
.040
.032
CR3212-5
NAS1739B5, NAS1739D5, NAS1769D4
CR3242-5
MS20426AD4
(Dimpled)
.063
NAS1739B4, NAS1739D4
MS20426AD5
(Countersunk)
.090
.080
.071
.063
.050
NAS1399B5, NAS1393D5
CR3212-5
NAS1739B5, NAS1739E5
NAS1769D5
CR3242-5
(See Note 1)
.063
.040
.032
NAS1739B6, NAS1739D6, NAS1769D6,
CR3212-6
CR3242-6
AN509-10 Screw with MS20365 Nut
.071
.090
NAS1739B5, NAS1739D5
NAS1739B6, NAS1739D6, CR3212-6
.071
.063
.032
NAS1769D6
CR3242-6
AN509-10 Screw with MS20365 Nut
.090
.032
NAS1739B6,NAS1739D6
AN509-10 Screw with MS20365 Nut
MS20426AD5
(Dimpled)
MS20426AD6
(Countersunk)
MS20426AD6
(Dimpled)
NOTE 1: Rework required. Countersink oversize to accommodate oversize rivet.
NOTE 2: Do not use blind rivets in high-vibration areas or to pull heavy sheets or extrusions together.
High-vibration areas include the nacelle or engine compartment including the firewall. Heavy
sheets or extrusions include spar caps.
18-6B
Revision 2
MODEL 210 & T210 SERIES SERVICE MANUAL
REPLACE
Fastener
Collar
* NAS178
NAS179
(See Note 1)
(See Note 1)
(See
(See
(See
(See
* NAS1054
Note 1)
Notes 1 and 2)
Note 1)
Note 1)
NAS179, NAS528
(See Note 2)
* NAS14XX
NAS1080C
NAS1080E
NAS1080G
* NAS529
NAS524A
WITH
DIAMETER
(See Note 3)
Fastener
Collar
* NAS1054
* NAS14XX
NAS179, NAS528
NAS1080C, NAS1080E,
NAS1080G
NAS524A
NAS1080C, NAS1080A6
NAS1080K
AN364, MS20364, MS21042
* NAS529
* NAS1446
* NAS7034
c NAS464
NAS1103
NAS1303
NAS6203
AN173
AN305, MS20305, MS21044,
MS21045
NAS14XX
NAS529
NAS1446
NAS7034
NAS464
NAS1103
NAS1305
NAS6203
NAS1080C, NAS1080E
NAS524A
NAS1080C, NAS1080A6
NAS1080K
AN364, MS20304, MS21042
* NAS529
* NAS1446
* NAS7034
NAS464
NAS1103
NAS1303
NAS6203
NAS524A
NAS1080C, NAS1080A6
NAS1080K
AN364, MS20364, MS21042
*
*
*
*
NAS1446
NAS1080C, NAS1080A6
NOTE 1: See appropriate tables for nominal diameters available.
NOTE 2: Available in oversize for repair of elongated holes. Ream holes to provide a
.001 inch interference fit.
NOTE 3: NAS1446 oversize only permitted as a replacement for NAS529.
* Steel shank fastener designed for drive-on collars.
* Steel shank fastener designed for squeeze-on collars. Installation requires sufficient space
for the tool and extended shank of the fastener.
Threaded fastener.
Revision 2
18-6C
MODEL 210 & T210 SERIES SERVICE MANUAL
V
WING
12 INCH WIDE HEAVY CANVAS
1 X 12 X 30-3/4
1 X 12 X 48
1 X 12 X 11
1 X 12 X 8
30.3/4
2 X 4 X 20
5 INCH COTTON WEBBING
42
2x 4
3/8 INCH DIAMETER
2
NOTE
BOLTS
X4
30
ALL DIMENSIONS ARE IN INCHES
Figure 18-1.
Wing and Fuselage Support Stands
18-66. BAFFLES. Baffles ordinarily require replacement if damaged or cracked. However, small
plate reinforcements riveted to the baffle will often
prove satisfactory both to the strength and cooling
requirements of the unit.
18-6D
Revision 3
18-67.
ENGINE COWLING.
18-68. REPAIR OF COWLING SKINS. If extensively
damaged, complete sections of cowling must be re-
MODEL 210 & T210 SERIES SERVICE MANUAL
c
GRIND
A
B
C
WING STATION
2.00
.75
2.00
2.00
40.50
25.50
26.50
205.00
ALL WING TWIST OCCURS BETWEEN STA. 26.50 AND STA. 205. 00.
CHECKING WING TWIST
If damage has occured 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 (42" minimum length of angle or equivalent), three
modified bolts and a protractor head with level.
1.
Check chart for applicable dimension for bolt length (A or B).
2.
Grind bolt shanks 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 first lateral row of rivets, aft of wing leading edge.
5.
Holding straightedge parallel to wing station, (staying as clear as possible from "cans"), place
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 outboard wing station, using dimensions specified in chart.
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.
Check
Forward or aft bolt
Figure 18-2. Checking Wing Twist
placed. Standard insert-type 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 Advisory
Circular 43.13-1 are also applicable to cowling.
Circular 43.13-1 are also applicable to cowling.
18-69. REPAIR OF REINFORCEMENT ANGLES.
Cowl reinforcement angles, if damaged, must be
replaced. Due to their small size they are easier
to replace than to repair.
18-70. REPAIR OF GLASS-FIBER CONSTRUCTED
COMPONENTS. Glass-fiber constructed components
on the aircraft may be repaired as stipulated in instructions furnished in SK182-12. 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 give better adhesion.
Revision 2
18-7
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
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 Catalog and installing
next to existing inboard weight the minimum length necesary for correct balance.
except that a length which contains at least two attaching screws must be used. If
necessary, lighten new weight or existing weights for correct balance.
10.
CENTERLINE ON BEAM MUST
BEAM ASSEMBLY
BE ALIGNED WITH CONTROL SURFACE
HINGE CENTERLINE
/
~HANGAR
ASSEMBLY
HINGE
CENTERLINE
CONTROL SURFACE
CHORD
ADD WASHERS AS NECESSARY
TO FINE BALANCE THE BEAM
ASSEMBLY
ADJUSTABLE
WEIGHT
HANGAR ASSEMBLY
(TO BE IN PROPER POSITION)
MANDREL
SLIDING
WEIGHT
READ CONTROL
SURFACE MOMENT
AT CENTER OF WEIGHT
BEAM ASSEMBLY
.
*MANDREL
Figure 18-3.
FLAT SURFACE
Cont'ol Surface Balancing (Sheet 2 of 5)
Revision 2
18-9
MODEL 210 & T210 SERIES SERVICE MANUAL
A balance in this range is "overbalance".
A balance in this range
is "underbalance".
Detail F
RUDDER
TRAILING EDGE
SPIRIT-LEVEL
PROTRACTOR
SLIDING
CENTER LINE
WEIGHT
CHORD
LINE
BALANCING
MANDREL
Detail H
LEVELED SURFACE
Figure 18-3.
18-10
Revision
2
HINGE POINT
Control Surface Balancing (Sheet 3 of 5)
ELEVATOR
MODEL 210 & T210 SERIES SERVICE MANUAL
AILERONS
*
MODEL 210 & T210 SERIES SERVICE MANUAL
CONTROL SURFACE BALANCE REQUIREMENTS
NOTE
Balance limits for control surfaces are expressed for
"Approved Flight" configuration. "Approved Flight"
configuration is that condition of the control surface as
prepared for flight of the airplane whether it be painted or
unpainted.
"Approved Flight" limits must never be exceeded when
the surface is in its final configuration for flight.
DEFINITIONS:
UNDERBALANCE is defined a the condition that exists when surface is trailing edge
heavy and I defined by a symbol (+). If the balance beam sliding weight must be-on the
leading edge side of the hinge line (to balane the control surface), the control surface is
considered to be underbalanced.
OVERBALANCE is defined as the condition that exists when surface is leading edge
heavy and is defined by a symbol (-). If the balance beam sliding weight must be on the
trailing edge side of the hinge line (to balance the control surface), the control surface is
conidered to be overbalanced.
CONTROL SURFACE
10
AILERON
4.25 to 11.16
RUDDER
-4.0 to 3.0
RIGHT ELEVATOR
0.0 to 12.1
LEFT ELEVATOR
0.0 to 12.1
Figure 18-3.
18-12
Revision
APPROVED FLIGHT CONFIGURATION
BALANCE LIMITS (Inch-Pounds)
Control Surface Balancing (Sheet 5 of 5)
MODEL 210 & T210 SERIES SERVICE MANUAL
B
-1/4
-B
-
-- '--'
.
PATCH
/- EXISTING SKIN
NOTE
For optimum appearance and
use flush rivets, dim-
I-/2airflow,
--
DOUBLER-
pled skin and patch and countersunk doubler.
SECTION THRU ASSEMBLED PATCH
A-A
V
'-':
, .:.
MARGIN = 2 X RIVET DIA.
-EDGE
L~
PATCH
R
2024-T3 ALCLAD
-
1/2" RADIUS
DA AGED AREA
EDWE MARGIN =
2 X RIVET DIA.
CLEAN OUT
,,
2 7 O
/6 X
RIVET DIA.
EDGE MARGIN
2XRIVET DIA.
=j
DOUBLER - 2024-TS3
ALCLAD
^21/2" RADIUS
_l
mREP AIR
PARTS
REPAIR
RADIUSR I
_LAR)
REPAIR PARTS IN CROSS SECTION
Figure 18-4. Skin Repair (Sheet 3 of 6)
18-14
.025
.032
.040
.051
1/8
1/8
1/8
5/32
MODEL 210 & T210 SERIES SERVICE MANUAL
FILLER -
2024-T4 ALCLAD
A-A
STRIP -
2024-T3 ALCLAD
1/4" EDGE MARGIN
CLEAN OUT
5 RIVETS EACH SIDE
OF DAMAGED AREA
ANGLE -
2024-T4 ALCLAD
STRINGER
PICK UP EXISTING SKIN RIVETS
MS20470AD4 RIVETS
A
SKIN
ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION
Figure 18-5.
Stringer and Channel Repair (Sheet 2 of 4)
18-19
MODEL 210 & T210 SERIES SERVICE MANUAL
MODEL 210 & T210 SERIES SERVICE MANUAL
FILLER-2024-T4 ALCLAD
DOUBLER-2024-T3 ALCLAD
3/4" RIVET
CLEAN OUT DAMAGED AREA
1/4" EDGE MARGIN
ANGLE-2024-T4 ALCLAD
RIB
ONE ROW RIVETS
AROUND DAMAGED
AREA
MS20470AD4 RIVETS
ORIGINAL PARTS
A-A
REPAIR PARTS
REPAIR PARTS IN
CROSS SECTION
Figure 18-6.
Rib Repair (Sheet 2 of 2)
18-23
MODEL 210 & T210 SERIES SERVICE MANUAL
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.
On cantilever wing, vertical size is limited by ability to install doubler clear of front fuel
spar or stringers outboard of spar. On flaps and ailerons, vertical size is limited by
ability to install doubler clear of front spar. (Also refer to figure 18-9.)
6.
Lateral size is limited to seven inches across trimmed out area.
7.
Number of repairs is limited to one in each bay. On cantilever wings, consider a bay in the
area forward of front fuel spar as if ribs extended to leading edge.
Use CR162-4
1" MAXIMUM RIVET
SPACING (TYPICAL)
DOUBLER NEED NOT--BE CUT OUT IF ALL
RIVETS ARE ACCESSIBLE
FOR BUCKING
/16" MINIMUM EDGE
MARGIN (TYPICAL)
//
.
/
TRIM OUT DAMAGED AREA
REPAIR DOUBLER
........ALCLAD
--
2024-T3
040" THICKNESS
FILLER MATERIAL
ORIGINAL PARTS
..-..
REPAIR PARTS
LEADING EDGE SKIN
LEADING EDGE SKIN
SAME THICKNESS
AS SKIN
Figure 18-8.
Leading Edge Repair
18-25
MODEL 210 & T210 SERIES SERVICE MANUAL
1/4" EDGE MARGIN
CLEAN OUT DAMAGED AREA
A-A
ANGLE -
2024-T4 ALCLAD
10 RIVETS EACH SIDE
OF DAMAGED AREA
FIREWALL ANGLE
FILLER
2024-T4 ALCLAD
A
MS20470AD4 RIVETS
FIREWALL
FUSELAGE SKIN
ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION
Figure 18-11.
18-28
Firewall Angle Repair
MODEL 210 & T210 SERIES SERVICE MANUAL
Use rivet pattern at wing station
25. 25 for repair from wing station
25. 25 to wing station 96. 00. Use
rivet pattern at wing station 110. 00
for lap splice patterns from wing
station 110.00 to 189. 00. See
figure 1-2 for wing stations.
-.
Use rivet spacing similar to the
pattern at wing station 110.00 at
leading edge ribs between lap
splices.\
Select number of flush rivets to be
used at each wing station leading
edge rib from table.
RIBS AND STRINGERS:
Blind rivets may be substituted for
solid rivets in proportionally
increased numbers in accordance
with the table.
SPARS:
Blind rivets may be installed in
wing spars only in those locations
where blind rivets were used during original manufacture, ie fuel
bay area of front spars on aircraft
with integral ruel bays.
NUMBER OF FLUSH RIVETS IN DIMPLED SKIN REQUIRED IN REPLACEMENT LEADING EDGE SKIN
WING
STATION
RIB
124
136
EXISTING
TACK RIVET
1
PATCH
SOLID
MS0426-4
155
-\
172
~
<
XISTING
18
15
22
18
10
12
11
10
89_____
BLIND
CR2248-4
13
12
EXISTING RIVET PATTERN
TYPICAL LEADING EDGE SECTION
Figure 18-12.
Bonded Leading Edge Repair
18-29/(18-30 blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
SECTION 19
EXTERIOR PAINTING
Page No.
TABLE OF CONTENTS
Aerofiche/Manual
MATERIALS LISTING .............
APPLICATION ....................
Interior Parts ..................
Exterior Parts Acrylic ...........
Exterior Parts Epoxy or
Polyurethane .................
MATERIALS LISTING .............
FACILITY .........................
3E21/19.1
3E22/19-2
3E22/19.2
3E22/19-2
3E22/19-2
3E23/19.3
3E24/19.4
APPLICATION ....................
Cleanup .......................
Prepriming ....................
Priming .......................
Prepainting ....................
Painting Overall -- White or Color
Stripes
Touchup ..................
Repair of Dents .................
3E24/19-4
3E24/19-4
3F1/19-5
3F/19-5
3F1/19-5
3F2/19-6
3F2/19-6
3F2/19-6
3F2/19-6
NOTE
Acrylic Lacquer is standard through serial
21061849
NOTE
This section contains a listing of standard factory materials and shows the area of their application. To
determine the paint number and color, refer to the aircraft trim plate and parts catalog. In all cases,
determine the type of paint because some types are not compatible with others. Contact Cessna Parts
Distribution (CPD 2) or a Cessna Service Station for materials acquisition information.
19-1.
MATERIALS LISTING.
MATERIAL
NO. /TYPE
PAINT
ACRYLIC
LACQUER
PRIMER
AREA OF APPLICATION
Used on exterior airframe.
ER-7 WITH
ER-4
ACTIVATOR
(THRU SERIAL 21061849)
Used with acrylic lacquer.
PRIMER
P60G2 WITH
R7K44
ACTIVATOR
Used with acrylic lacquer.
THINNER
T-8402A
Used to thin acrylic lacquer and for burndown.
SOLVENT
#2 SOLVENT
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 paint from aircraft, do not allow
stripper to contact ABS parts.
Revision 3
19-1
MODEL 210 & T210 SERIES SERVICE MANUAL
19-2.
APPLICATION
9-3. INTERIOR PARTS (Finish Coat of Lacquer)
a. Painting of Spare Parts.
1. Insure a clean surface by wiping with Form
Tech AC 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 Form
Tech AC to remove surface contamination.
CAUTION
CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
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
or 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 Form
Tech AC to remove surface contamination.
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.
19-5.
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
Lacquer paints can be successfully spotted in.
19-4. 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 Form
Tech AC 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. 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
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 Form
Tech AC to remove surface contamination.
3AToN
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
3. Apply a primer compatible with Epoxy or
Polyurethane topcoat.
4. After the part is thoroughly dry it is ready
for the topcoat.
NOTE
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.
MODEL 210 & T210 SERIES SERVICE MANUAL
19-6.
MATERIALS LISTING.
NOTE
Enflex III is standard beginning 21061850
thru 21062000 and 21062002 thru 21062009,
21062011, 21062012, 21062019, 21062023
thru 21062025, 21062027 thru 21062029,
2106231 thru 2106233, 21062035, 21062037
thru 21062039, 21062043, 21062044, 21062046
thru 21062049, 21062054, 21062055, 21062057,
21062059, 21062065, 21062069. and 21062072.
ENMAR MODIFIED URETHANE
MATERIAL
NO/TYPE
AREA OF APPLICATION
ENFLEX III ENAMEL
Standard Exterior, and Stripe Only configuration
ENFLEX III ADDUCT
Catalyst for Enflex IIIEnamel
ACCELERATOR
URETHANE ACCELERATOR
120-975
Used to speed curing on stripes
PRIMER
WASH PRIMER EX-ER-7
Used to prime aircraft for Enflex IIItopcoat
REDUCER
T-ER-4
Used to thin EX-ER-7
THINNER
Jet Glo
86T-10399 (110-655)
Used to thin Enflex III
110-805
Used to thin Enflex III
110-996
Used to slow curing time
PAINT
RETARDER
NOTE
Imron is Standard beginning with 21062001,
21062010, 21062013 thru 21062018, 21062020
thru 21062022, 21062026, 21062030, 21062034,
21062036, 21062040 thru 21062042, 21062045,
21062050 thru 21062053, 21062056, 21062058,
21062060 thru 21062064, 21062066 thru
21062068, 21062070, 21062071, 21062073
and all 1978 Models.
IMRON MODIFIED URETHANE
MATERIAL
PAINT
NO/TYPE
AREA OF APPLICATION
IMRON ENAMEL
Used as corrosion proof topcoat
IMRON 192S Activator
Catalyst for Imron Enamel
PRIMER
WASH PRIMER P60G2
Used to prime aircraft for Imron Enamel
REDUCER,
THINNER
IMRON Y8485S Reducer
Used to thin Imron Enamel
Catalyst Reducer R7K44
Used to reduce P60G2
NOTE
Do not paint pitot tube. gas caps, or aileron gap seals.
Also
do not paint antenna covers which were not painted at the factory.
19-3
MODEL 210 & T210 SERIES SERVICE MANUAL
MATERIAL
NO/TYPE
AREA OF APPLICATION
STRIPPER
Strypeeze Stripper
Used to strip primer overspray
CLEANER
Form Tech AC
Used to clean aircraft exterior and to remove
grease, bug stains, etc.
Klad Polish
Used to clean aluminum finish
808 Polishing Compound
Used to rub out overspray
SOLVENT
(MEK) Methyl Ethyl Ketone
Used to tack aircraft prior to topcoat
CLOTH
HEX Wiping Cloth
Used with solvent to clean aircraft exterior
FILLER
White Streak
Used to fill small dents
MASKING
Class A Solvent Proof Paper
Used to mask areas not to be painted
Tape Y218
Used for masking small areas
Tape Y231
Used for masking small areas
19-7. FACILITY. Painting facilities must include
the ability to maintain environmental control; temperature at 65ºF., and a positive pressure inside to
preclude the possibility of foreign material damage.
All paint equipment must be clean, and accurate
measuring containers available for mixing protective
coatings. Modified Urethane has a pot life of four to
eight hours, depending on ambient temperature and
relative humidity. Use of approved respirators while
painting is a must, for personal safety. All solvent
containers should be grounded to prevent static buildup. Catalyst materials are toxic, therefore, breathing fumes or allowing contact with skin can cause
serious irritation. Material stock should be rotated
to allow use of older materials first, because its
useful life is limited. All supplies should be stored
in an area where temperature is higher than 50 F.,
but lower than 90 F. Storage at 90 F is allowable
to roomfor
temperature
mixing and use.spray
Modified urethane paint requires a minimum of seven
temperature is lower, curing time will be extended a
maximum of 14 days. During the curing period, in-
discriminate
use of
discriminate use
of masking
masking tape,
tape, abrasive
abrasive polishes,
polishes,
or cleaners can cause damage to finish. Desirable
temperature for modified urethane is 60ºF.
curing
curing
temperature
formodfied urethane
is
60F.
for a resulting satisfactory finish.
b. Wipe excess sealer from around windows and
skin laps, using Form Tech AC. Mask windows,
ABS parts and any other areas not to be primed,
with 3M tape and Class A Solvent-Proof Paper.
Care must be exercised to avoid cuts, scratches or
gouges by metal objects to all plexiglass surfaces,
because cuts and scratches may contribute to crazing
and failure of plexiglass windows.
c. Methyl Ethyl Ketone (MEK) solvent should be
used for final cleaning of airplanes prior to painting.
The wiping cloths shall be contaminant and lint free
HEX. Saturate cloth in the solvent and wring out so
it does not drip. Wipe the airplane surface with the
solvent saturated cloth in one hand, and immediately
dry with a clean cloth in the other hand. It is important to wipe dry solvent before t evaporates. Avoid
contact of MEK with plexiglass, as crazing will result.
When an airplane has paint or zinc chromate overon the exterior. stripper may be used to re-
move the overspray.
The stripped may be applied
by brush and will require a ew minutes to soften the
overspray. Heavy coatings may require more than
one application of the stripper. Use extreme care to
prevent stripper from running into faying surfaces on
corrosion proofed airplanes After removal of the
removal of the
corrosion proofed airplanes. After
overspray, clean the airplane with Methyl Ethyl Ketone (MEK) solvent in the prescribed manner.
NOTE
19-8.
APPLICATION.
19-9. CLEAN UP.
a. Inspect airplane for any surface defects, such as
dents or unsatisfactory previous repairs, and correct
according to Paragraph 18-9.
It is imperative that clean solvent be used in
cleaning airplanes. Dispose of contaminated
solvent immediately. Fresh solvent should
be used on each airplane.
WARNING
Use explosion proof containers for storing
wash solvents and other flammable materials.
19-4
MODEL 210 & T210 SERIES SERVICE MANUAL
19-10.
PREPRIMING.
NOTE
Enflex III is standard beginning 21061850
thru 21062000 and 21062002 thru 21062009,
21062011, 21062012, 21062019, 21062023
thru 21062025, 21062027 thru 21062029,
2106231 thru 2106233, 21062035, 21062037
thru 21062039, 21062043, 21062044, 21062046
thru 21062049, 21062054, 21062055, 21062057,
21062059, 21062065, 21062069, and 21062072.
a. Above serialized aircraft have Enmar Wash
Primer EX-ER-7, Enflex IIIEnamel for overall
color and stripes.
b. Mix one to one, EX-ER-7 primer with T-ER-4
Reducer by volume. Mix only in stainless steel or
lined containers only. After mixing allow primer to
set for 30 minutes before spraying. Pot life of the
mixed primer is six (6) hours. All mixed material
should be discarded if not used within this time.
Pot pressure during spraying should be approximately 10 PSI ± 1 PSI. Air pressure should be 40 to 50
PSI at the gun. Blow loose contaminant of the aircraft with clean, dry air. Check all tapes to make
sure it adheres properly. Cover the flap tracks,
nose gear strut tube, wheels, and shimmy dampener
rod ends. ABS parts and other preprimed parts do
not receive wash primer.
NOTE
19-11. PRIMING.
a. Apply primer in one wet even coat. Dry film
thickness to be ,0003 to ,0005 inches. Do not topcoat until sufficiently cured. When scratching with
firm pressure of the fingernail does not penetrate
the coating, the primer is cured. Primer should be
topcoated within four hours after application.
19-12.
PREPAINTING
NOTE
Enflex IIIis standard beginning 21061850
thru 21062000 and 21062002 thru 21062009,
21062011, 21062012, 21062019, 21062023
thru 21062025, 21062027 thru 21062029,
2106231 thru 2106233, 21062035, 21062037
thru 21062039, 21062043, 21062044, 21062046
thru 21062049, 21062054, 21062055, 21062057,
21062059, 21062065, 21062069, and 21062072.
a. On above serialized aircraft, mix the required
amount of Enflex IIIwith Enflex IIIAdduct in a 4 to 1
ratio by volume. Mix thoroughly, and allow to stand
for approximately 30 minutes before spraying.
Enflex IIIcan be thinned with Jet Glo thinner 86T10399 (110-655) to obtain spraying viscosity, which
should be checked after four hours and adjusted if
necessary.
NOTE
Imron is Standard beginning with 21062001,
21062010, 21062013 thru 21062018, 21062020
thru 21062022, 21062026, 21062030, 21062034,
21062036, 21062040 thru 21062042, 21062045,
21062050 thru 21062053, 21062056, 21062058,
21062060 thru 21062064, 21062066 thru
21062068, 21062070, 21062071, 21062073
and all 1978 Models.
Imron is Standard beginning with 21062001,
21062010, 21062013 thru 21062018, 21062020
thru 21062022, 21062026, 21062030, 21062034,
21062036, 21062040 thru 21062042, 21062045,
21062050 thru 21062053, 21062056, 21062058,
21062060 thru 21062064, 21062066 thru
21062068, 21062070, 21062071, 21062073
and all 1978 Models.
c. Corrosion proofed and standard aircraft will
receive Sherwin Williams Primer P60G2, DuPont
Imron Enamel for over all color, and for stripes,
d. Mix 1 part P60G2 primer with 1 1/2 parts
R7K44 catalyst reducer, by volume. Mix in stainless steel or lined containers only. After mixing
allow primer to set for 30 minutes before spraying.
Pot life of the mixed primer is six (6) hours, all
mixed materials should be discarded if not used
within that time limit. Pot pressure during spraying
should be approximately 10 PSI ± 1 PSI. Air pressure
should be 40 to 50 PSI at the gun. Blow loose contaminant off the airplane with clean, dry air. Check
all tapes to make sure they adhere properly. Cover
the flap tracks, nose gear strut tube, wheels, and
shimmy dampener rod ends. ABS parts and other
preprimed parts do not receive wash primer.
b. On standard aircraft mix the required amount of
Imron with Imron 192S Activator in a 3 to 1 ratio by
volume. Mix thoroughly, and begin spraying immediately, because there is no induction time requirement
Imron can be thinned to spraying viscosity with Y8485S
Imron Reducer. Viscosity should be checked and adjusted after four hours if necessary.
WARNING
AIRCRAFT SHOULD BE GROUNDED
PRIOR TO PAINTING TO PREVENT STATIC
ELECTRICITY BUILDUP AND DISCHARGE.
c. When applying modified urethane finishes, the
painter should wear an approved respirator, which
has a dust filter and organic vapor cartridge, or an
air supplied respirator. All modified urethane finishes contain some isocyanate, which may cause irritation to the respiratory tract or an allergic reaction.
Individuals may become sensitized to isocyanates.
d. The pot life of the mixture is approximately 6-8
hours at 75°F (24°C). Pot pressure should be
approximately 12 PSI during application. Air pressure
at the gun should be 40 to 50 PSI.
e. Scuff sand the primer only where runs or dirt
particles are evident. Minor roughness or grit may
be removed by rubbing the surface with brown Kraft
Revision 3
19-5
MODEL 210 & T210 SERIES SERVICE MANUAL
paper which has been thoroughly wrinkled. Unmask
ABS and other preprimed parts and check tapes.
Clean surface with a jet of low pressure-dry air.
f. Painting of the stripe should be done with two or
three wet, even coats. Dry coats will not reflow and will
leave a grainy appearance. Stripes may be force dried or
air dried. Film thickness of a stripe is approximately 1.0
19-13. PAINTING OVERALL -- WHITE OR COLOR.
mil.
a. Complete painting of the plane should be done with
two or three wet, even coats. Dry coats will not reflow
and will leave a grainy appearance.
b. Allow a five minute period for the finish to flash off
before moving aircraft to the oven.
c. Move to the force dry oven and dry for approximately 1 ½ hours at 120°F to 140°F (49°C to 60°C).
d. Dry film thickness of the overall color should be
between 1.3 and 2.0 mils. Films in excess of 3.0 mils are
not desirable.
g. Do not remove masking tape and paper until the
paint has dried to a "dry to touch" condition. Care
should be exercised in removal of the masking to prevent damage to the finish.
h. Uncured urethane finishes are sensitive to moisture, therefore, should be stored out of rain until
cured.
19-14. STRIPES.
a. Remove airplane from the oven. Allow airplane
to cool to room temperature before masking.
b. Mask stripe area using 3M Tape Y231 or 3M Tape
Y218 and Class A solvent proof paper. Double tape all
skin laps to prevent blow by.
c. Airplanes which will have a stripe only configuration shall be masked, cleaned, and primed, in stripe
area only.
d. If the base coat is not over 72 hours old, the
stripe area does not require sanding. If sanding is
necessary because of age or to remove surface defects, use #400 or #600 sand paper. Course paper
will leave sand marks which will decrease gloss and
depth of gloss of the finish. The use of power sanders
should be held to a minimum; if used, exercise care
to preclude sanding through the white base coat. Wipe
surface to be striped with a tack cloth and check all
tapes.
e. Stripe colors on Enflex III,Jet Glo, or acrylic base
coat will be Acry Glo, and on Imron modified urethane
base coat will be Imron Enamel. When mixing tints for
stripes, stir the containers for at least 20 minutes before
weighing out the required masses. Mix Acry Glo using
three volumes of 571 Series Base with one volume of
581-091 catalyst; thin mixture with 110-701 or 110-755
thinners 20% to 25% by volume (18 to 25 seconds in a
No. 2 Zahn cup). Mix Imron using eight volumes of base
with one volume VG-Y-1421 catalyst (ratio three to one
if 1925 activator used); thin with Cessna Thinner No. 1
(18 to 20 seconds in a No. 2 Zahn cup).
19-6
Revision 3
19-15. TOUCHUP.
When necessary to touch up or refinish an area. the
defect should be sanded with 1400 and followed by 2600
sand paper. Avoid, if possible, sanding through the
primer. If the primer is penetrated over an area 1/2
inch square or larger, repriming is necessary. Avoid
spraying primer on the adjacent paint as much as possible. Since urethane finishes cannot be "spotted in"
repairs should be in sections extending to skin laps or
stripe lines.
a. Dry overspray and rough areas may be compounded out with DuPont #808 rubbing compound.
b. Grease, bug stains, etc., may be removed from
painted surfaces with Form Tech AC. Klad Polish
may be used on bare aluminum to remove stains, oxides, etc.
c. Rework areas, where paint or primer removal is
required, may be stripped with Strypeeze Paint Removal. All traces of stripper must be removed before
refinishing.
19-16. REPAIR OF DENTS.
a. To repair dents use White Streak Filler or equivalent. Mix White Streak. in the correct proportion as
recommended by the manufacturer.
b. Do not apply White Streak Filler over paint. All
paint shall be removed in the repair area and the aluminum surface sanded lightly to increase adhesion.
Apply the White Streak to a level slightly above the
surrounding skin. After drying for 10 - 15 minutes,
sand the filler flush with the skin surface, using care
to feather the edges.
MODEL 210 &T210 SERIES SERVICE MANUAL
SECTION 20
WIRING DIAGRAMS
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
Circuit Function and Specific
Circuit Code Letters ...............
Circuit Function and Wires ..........
Wiring Diagram Serial Numbers
Vs. Aircraft Serial Numbers ....
D.C. POWER
Battery Circuit .................
Ground Service Receptacle (OPT)
Alternator System ..............
Circuit Breakers ...............
Alternator System ..............
*
Alternator System, 95 Amp (OPT)
Circuit Breakers ...............
Alternator System, 60 Amp ......
Alternator System, 95 Amp (OPT)
Circuit Breakers ...............
Battery Circuit .................
Circuit Breakers ...............
Circuit Breakers ...............
Circuit Breakers ...............
Standby Generator (OPT) .......
Ground Service (OPT) ...........
Battery Circuit .................
Alternator System, 60 Amp ......
Standby Generator (OPT) .......
Dual Alternator ................
Dual Alternator ................
Dual Alternator ................
Dual Alternator ................
Volt-Ammeter (OPT) ...........
Battery Circuit - OPT (With Dual
Alternators) ..................
Alternator System, 95 Amp (OPT)
Circuit Breakers ...............
Circuit Breakers ...............
Circuit Breakers ...............
Volt-Ammeter (OPT) ...........
IGNITION
Magneto
....................
ENGINE CONTROLS
Starter System .................
Starter System .................
FUEL OIL
Fuel Pump System
.............
ENGINE INSTRUMENTS
Cylinder Head Temp ............
Hourmeter (OPT) ...............
Hourmeter (OPT) ...............
Fuel Gage System ..............
Instrument Cluster .............
Fuel Gages.
..............
Fuel Gages .
..............
Oil Temp & Cylinder Head
Temperature .................
Oil Temp & Cylinder Head
Temperature .................
Fuel Gages ....................
Fuel Gages ....................
Fuel Totalizer/Clock (OPT) ......
Oil Temp & Cylinder Head
Temperature .................
FLIGHT INSTRUMENTS
Turn Indicator .................
Encoding Altimeter .............
OTHER INSTRUMENTS
3H22/20-62
Clock ..........................
Ammeter ...................... 3H23/20-63
3F7/20-2A
Ammeter
..........
3H24/20-64
3F8/20-3
Digital Clock (OPT) ............. 311/20-65
|
Ammeter ...................... 312/20-66
3F8/20-3
LIGHTING
Electroluminescent Panel .......
313/20-67
3F10/20-5
Electroluminescent Panel ....... 314/20-68
3F11/20-6
Instrument Lights .............. 315/20-69
3F12/20-7
Instrument Lights ...........
317/20-71
3F14/20-9
Lights
.............
318/20-72
3F15/20-10
NavigationLights .............. 319/20-73
3F16/20-11
Dome, Courtesy & Baggage Lights 3I11/20-75
3F17/20-12
Console & Compass Lights ....... 3113/20-77
3F19/20-14
Eyebrow Lights ................. 3115/20-79
3F21/20-16
Wing Tip Strobe Lights (OPT) ....
3117/20-81
3F23/20-18
Wing Tip Strobe Lights (OPT) ....
3119/20-83
3F24/20-19
Flashing Beacon Light .......... 3120/20-84
3G1/20-20
Flashing Beacon Light .......... 3122/20-86
3G3/20-22
Post Lighting (OPT) ............. 3123/20-87
3G4/20-23
Map Light, Control Wheel .......
3J1/20-89
3G5/20-24
Instrument Lights.
.......... 3J2/20-90
3G7/20-26
Instrument Lights.
.......... 3J4/20-92
3G8/20-27
Electroluminescent Panel (OPT) . 3J5/20-93
3G9/20-28
Electroluminescent Panel (OPT) . 3J6/20-94
3G10/20-29
Post Lighting (OPT) ...........
3J7/20-95
3G11/20-30
Post Lighting (OPT) ...........
3J8/20-96
3G13/20-32
Post Lighting (OPT) ...........
3J9/20-97
3G14/20-32A
Map Light, Control Wheel ....... 3J10/20-98
3G15/20-33
Flood, Engine Instr, & Radio
3G16/20-34
Dial Lights ................... 3J11/20-99
Console & Compass Lights ....... 3J12/20-100
3G17/20-35
Console & Compass Lights ....... 3J14/20-102
3G19/20-37
Vertical Tail Illumination
3G20/20-38
Light (OPT) ................... 3J15/20-103
3G21/20-38A
Flood, Engine Instr, & Radio
3G22/20-38BDial Lights ................... 3J16/20-104
3G23/20-39
Map Ligt, Control Wheel ....... 3J17/20-105
LANDING GEAR
3G24/20-40
Landing Gear Control System ....
3J18/20-106
Landing Gear Control System
.
3J19/20-107
3H1/20-41
Landing Gear Control System ....
3J21/20-109
3H2/20-42
Landing Gear Control System ....
3J22/20-110
Landing Gear Control System
3J24/20-112
3H3/20-43
Landing Gear Control System ....
3K1/20-113
Landing Gear Control System .... 3K2/20-114
3H5/20-45
Landing Gear Control System ....
3K3/20-115
3H6/20-46
Landing Gear Control System ...
3K4/20-116
3H7/20-47
HEATING, VENTILATING, & DE-ICING
3H8/20-48
Cigar Lighter .................. 3K6/20-118
3H9/20-49
Heated Pitot Tube and Heated
31111/20-51
Stall Warning System .........
3K7/20-119
3H13/20-53
Light, Ice Detector (OPT) ........
3K8/20-120
Wing De-cing System (OPT)
3K9/20-121
3H14/20-54
Windshield Anti-Ice System (OPT) 3K10/20-122
Prop De-Icing System 3 Blade (OPT) 3K11/20-123
3H15/20-55
Cigar Lighter .................. 3K13/20-125
3H16/20-56
Wing &Stabilizer De-Icing
3H17/20-57
System (OPT) ................. 3K14/20-126
31118/20-58
Heated Pitot Tube & Heated
Stall Warning System |
3H19/20-59
Known Icing (OPT) ............ 3K15/20-127
Heated Pitot Tube & Heated
3H20/20-60
Stall Warning System 3H21/20-61
Known Icing (OPT) ............
3K16/20-128
Revision 3
20-1
MODEL 210 &T210 SERIES SERVICE MANUAL
TABLE OF CONTENTS
Page No.
Aerofiche/Manual
Heated Pitot & Heated Stall
Warnin System - Known
Icing(OPT) .................
Windshield Anti-Ice System
Known Icing (OPT) ............
Wing & Stabilizer De-Ice System 3 Cycle (OPT) .................
Wing & Stabilizer De-Ice System 3 Cycle (OPT) .................
Air Conditioner (OPT) ..........
Air Conditioner (OPT) ..........
Heated Pitot Tube & Heated
Stall Warning System .........
20-2
Revision 3
3K17/20-129
3K18/20-130
3K19/20-131
3K20/20-132
3K22/20-133
3K23/20-134
3K24/20-135
CONTROL SURFACES
Wing Flaps .....................
Electric Elevator Trim (OPT) ....
Electric Elevator Trim (OPT) ....
Electric Elevator Trim (OPT) ....
Electric Elevator Trim (OPT) ....
Electric Elevator Trim (OPT) ....
Wing Flaps .....................
WARNING AND EMERGENCY
Duel WarningUnit .............
Dual Warning System ...........
Dual Warning System ...........
Duel Warning Unit .............
Duel Warning Unit .............
Duel Warning Unit.
Low Vacuum Warning Light (OPT)
3L120-136
3L13/20-138
3L5/20-140
3L7/20-142
3L9/20-144
3L10/20-145
3L12/20-147
3L13/20-148
3L15/20-150
3L17/20-152
3L18/20-153
320/20-155
3L21/20-156
3L22/20-157
MODEL 210 & T210 SERIES SERVICE MANUAL
CIRCUIT FUNCTION AND SPECIFIC CIRCUIT CODE LETTERS
A - Armament
B - Photographic
C - Control Surface
CA - Automatic Pilot
CC - Wing Flaps
CD - Elevator Trim
D - Instrument (Other Than Flight or Engine
Instrument)
DA - Ammeter
DB - Flap Position Indicator
DC - Clock
DD - Voltmeter
DE - Outside Air Temperature
DF - Flight Hour Meter
E - Engine Instrument
EA - Carburetor Air Temperature
EB - Fuel Quantity Gage and Transmitter
EC - Cylinder Head Temperature
ED - Oil Pressure
EE - Oil Temperature
EF - Fuel Pressure
EG - Tachometer
EH - Torque Indicator
EJ - Instrument Cluster
F - Flight Instrument
FA - Bank and Turn
FB - Pitot Static Tube Heater and Stall Warning
Heater
- FC - Stall Warning
FD - Speed Control System
FE - Indicator Lights
G - Landing Gear
GA - Actuator
GB - Retraction
GC - Warning Device (Horn)
GD - Light Switches
GE - Indicator Lights
H - Heating, Ventilating and De-Icing
HA - Anti-icing
HB - Cabin Heater
HC - Cigar Lighter
HD - De-ice
HE - Air Conditioners
HF - Cabin Ventilation
J - Ignition
JA - Magneto
K - Engine Control
KA - Starter Control
KB - Propeller Synchronizer
L - Lighting
LA - Cabin
LB - Instrument
LC- Landing
LD - Navigation
L - Taxi
LF - Rotating Beacon
LG - Radio
LH - De-ice
LJ - Fuel Selector
LK - Tail Floodlight
M - Miscellaneous
MA - Cowl Flaps
MB - Electrically Operated Seats
MC - Smoke Generator
MD - Spray Equipment
ME - Cabin Pressurization Equipment
MF - Chem 02 - Indicator
P - D. C. Power
PA - Battery Circuit
PB - Generator Circuits
PC - External Power Source
Q - Fuel and Oil
QA - Auxilliary Fuel Pump
QB - Oil Dilution
QC - Engine Primer
QD - Main Fuel Pumps
QE - Fuel Valves
R - Radio (Navigation and Communication)
RA - Instrument Landing
RB - Command
RC - Radio Direction Finding
RD - VHF
RE - Homing
RF - Marker Beacon
RG - Navigation
RH - High Frequency
RJ - Interphone
RK- UHF
RL - Low Frequency
RM- Frequency Modulation
RP - Audio System and Audio Amplifier
RR - Distance-Measuring Equipment (DME) RS - Airborne Public Address System
S - Radar
U - Miscellaneous Electronic
UA - Identification - Friend or Foe
W - Warning and Emergency
WA - Flare Release
WB - Chip Detector WC - Fire Detection System
X - A.C. Power
Revision 2
20-2A/(20-2B blank)
MODEL 210 & T210 SERIES SERVICE MANUAL
BASE
MODEL 210 & T210 SERIES SERVICE MANUAL
CROSS REFERENCE LISTING OF SERIAL REQUEST NUMBERS
LISTED ON DIAGRAMS VS. AIRCRAFT SERIAL NUMBERS (CONT).
SR No.
AIRCRAFT SERIAL No.
SR No.
AIRCRAFT SERIAL No.
SR8153
21060719
SR9310
T21063641, *P21000386
SR8259
21061041
*SR9361
SR8297
21061103
SR9384
21062955
SR8394
21061315
SR9427
21062969, *P21000120
SR8426
21061296
SR9429
21063299, *P21000257
SR8464
21062274
SR9465
21063369, *P21000279
6R8465
P21000001 thru P21000150
SR9556
21063953, T21067300 & *P21000405
SR8482
21061230
SR9583
21063547, P21000344
SR8499
21061574
SR9634
21064136
SR8552
21061617
*SR9635
SR8633
21061984
SR9711
21064064, *P21000535
SR8656
21061627
SR9742
21064038, *P21000553
SR8784
21062954
SR9785
21064136. *P21000591
SR8785
SR8861
SR8863
P21000151
P21000591
P21000151
SR9953
21064536
21062274, *P21000001
*SR9954
P21000761
SR10056
21064536,
SR10122
21064536, *P21000761
SR10250
21064559, *P21000771
SR10061
21064198
2104198
21064773
21062274 thru 21062953
*P21000001 thru P21000150
SR8938
SR.8938 21062250
21062250
SR8970
21062250, *P21000001
SR9014
21062727, *P21000041001
SR10101
P21000761
MODEL 210 & T210 SERIES SERVICE MANUAL
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MODEL 210 & T210 SERIES SERVICE MANUAL
T
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MODEL 210 & T210 SERIES SERVICE MANUAL
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MODEL 210 & T210 SERIES SERVICE MANUAL
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Source Exif Data:
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