D2006 4 13 182 AND SKYLANE SERIES (1969 THRU 1976) Cessna_182_Skylane_1969 1976_D2006 Cessna 1969 1976

User Manual: Cessna_182_Skylane_1969-1976_D2006-4-13

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REVISION
MODEL
182
AND SKYLANE SERIES
1969 THRU 1976

SERVICE MANUAL
REVISION 4
1 MARCH 2004

D2006R4-1 3
INSERT THE FOLLOWING REVISED
PAGES INTO THE BASIC MANUAL

Cessna

ATextron Company

Service Manual
1969 THRU 1976
MODEL 182
AND SKYLANE SERIES

Member of GAMA
FAA APPROVAL HAS BEEN OBTAINED ON TECHNICAL DATA IN THIS PUBLICATION THAT AFFECTS
AIRPLANE TYPE DESIGN.

REVISION 4 TO THE BASIC MANUAL IS BEING SUPPLIED TO PROVIDE ADDITIONAL
INFORMATION NECESSARY TO MAINTAIN THE AIRPLANE AND INCORPORATES TEMPORARY
CHANGE 1 DATED 5 SEPTEMBER 1977, TEMPORARY REVISION 1 DATED 3 OCTOBER 1994,
TEMPORARY REVISION 2 DATED 7 JANUARY 2000, AND TEMPORARY REVISION 3 DATED
7 OCTOBER 2002.

15 SEPTEMBER 1972

REVISION 4

1 March 2004

CESSNA AIRCRAFT COMPANY

MODEL 182 SKYHAWK SERIES
SERVICE MANUAL

LIST OF EFFECTIVITY PAGES
INSERT THE LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.
Dates of issue for original and revisions are:
Change ......... 3........ 1 October 1975

Original..........0 ...... 15 September 1972
Change..........1 ......

1 November 1973

Change..........2 ......

1 September 1974

Revision ........ 4........ 1 March 2004

Note: The portion of the text affected by the revision is indicated
by a vertical line in the outer margin of the page.
*The asterisk indicates pages revised, added, or deleted by current revision.
Page
No.

Revision
No.

* Title ............................. 4
*A thru B ...................... 4
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Revision 4
Mar 1/2004

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A
© Cessna Aircraft Company

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
LIST OF EFFECTIVITY PAGES (CONT.)
Page
No.

Revision
No.

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...... 3
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....
... 3
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........... 3
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..............
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Revision 4
Mar 1/2004

TABLE OF CONTENTS
Page

SECTION
1

GENERAL DESCRIPTION . . . . ..

. ..

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

GROUND HANDLING, SERVICING, CLEANING,
INSPECTION ........

1-1........

L UBRICATION AND
2-1
3-1

FUSELAGE

WINGS AND EMPENNAGE ...........

4-1

LANDING GEAR AND BRAKES ....

5-1

AILERON CONTROL SYSTEM ....

6-1

WING FLAP CONTROL SYSTEM .......

7-1

ELEVATOR CONTROL SYSTEM ....

8-1

ELEVATOR TRIM TAB CONTROL SYSTEM . . .

9-1

1.0

RUDDER AND RUDDER TRIM CONTROL SYSTEM

10-1

ENGINE ..........

11-1

1.2

FUEL SYSTEM ........

12-1

113

PROPELLER AND GOVERNOR ....

13-1

114

UTILITY SYSTEMS .......

14-1

1.5

INSTRUMENTS AND INSTRUMENT SYSTEMS . .

15-1

ELECTRICAL SYSTEMS ......

16-1

ELECTRONIC SYSTEMS (DELETED) (See Page iii)

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

18-1

PAINTING ..........

19-1

WIRING DIAGRAMS...............

20-1

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

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
INTRODUCTION

General.
WARNING:

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, 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.

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
Newsletters, 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 information available at Cessna Authorized Service Stations or through the Cessna
Propeller Aircraft 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 are communicated
to the field through Cessna's Authorized Service Stations and/or through Cessna's subscription
services.

Inspection, maintenance and parts requirements for 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 installation.

REVISIONS, REISSUES, and TEMPORARY REVISIONS can be purchased from your Cessna
Service Station or directly from Cessna Propeller Aircraft Product Support, Department 751, Cessna
Aircraft Company, P.O. Box 7706, Wichita, Kansas 67277-7706.

This manual contains factory recommended procedures and instructions for ground handling,
servicing and maintaining Cessna Model 182-Series and F182-Series aircraft. This includes the
Model A182, which is manufactured by Fuerza Aerea Argentina, Area de Material, Cordoba.

All supplemental service information concerning this manual is supplied to all appropriate Cessna
Service Stations so they have the latest authoritative recommendations for servicing these Cessna
airplanes. It is recommended that Cessna owners utilize the knowledge and experience of the
Cessna Service Station.

Revision 4
Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL

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

SKYLANE ONLY
REIMS

ARGENTINE 182

AMC 182

MODEL
182M
182N
182N
182P
182P
182P
182P
182P
F182P

BEGINNING
SERIAL NUMBER
18259306
18260056
18260446
18260826
18261426
18262466
18263476
18264296
F1820001

1969
1970
1971
1972
1973
1974
1975
1976

A182M
A182N
A182N
A182N
A182N
A182N
A182N
A182N

NONE
A182-0117
NONE
NONE
A182-0137
NONE
NONE
NONE

ENDING
SERIAL NUMBER
18260055
18260445
18260825
18261425
18262465
18263475
18264295
18265175
F18200025

A182-0136

A182-0146

Coverage and Format.
A.

The Cessna Model 182-Series Service Manual has been prepared to help maintenance
personnel in servicing and maintaining the Model 182-Series. This manual provides the
necessary information required to enable the mechanic to service, inspect, troubleshoot,
remove and replace components or repair systems.

Technical Publications are also available for the various components and systems which are
not covered in this manual. These manuals must be utilized as required for maintenance of
those components and systems, and may be purchased from the manufacturer.

Temporary Revisions.
A.

MODEL YEAR
1969
1970
1971
1972
1973
1974
1975
1976
1976

Additional information which becomes available may be provided by temporary revision. This
service is used to provide, without delay, new information which will assist in maintaining safe
flight/ground operations. Temporary revisions are numbered consecutively. Temporary
revisions are normally incorporated into the maintenance manual at the next regularly
scheduled revision.

Material Presentation.
A.

Revision 4
Mar 1/2004

This Service Manual is available on paper, aerofische or Compact Disc (CD/ROM); The CD
ROM contains the Service Manuals, Illustrated Parts Catalogs and Avionics Manuals.
iii/ (iv Blank)
© Cessna Aircraft Company

FOREWORD
This manual contains factory recommended procedures and instructions for ground handling, servicing and maintaining Cessna
Model 182-Series aircraft. This includes the Model A182, which
is manufactured by Fuerza Aerea Argentina, Area de Material,
Cordoba.
Besides serving as a reference for the experienced mechanic,
this book also covers step-by-step procedures for the less experienced man. This manual should be kept in a handy place for ready
reference. If properly used, it will better enable the mechanic to
maintain Cessna Model 182-Series aircraft and thereby establish a
reputation for reliable service.
The information in this book is based on data available at the
time of publication and is supplemented and kept current by service
letters and service news letters published by Cessna Aircraft Company. These are sent to all Cessna Dealers so that they have the
latest authoritative recommendations for servicing Cessna aircraft.
Therefore, it is recommended that Cessna owners utilize the knowledge and experience of the factory-trained Dealer Service Organization.
In addition to the information in this Service Manual, a group
of vendor publications is available from the Cessna Service Parts
Center which describe complete disassembly, overhaul and parts
breakdown of some of the various vendor equipment items. A listing of the available publications is issued periodically in service
letters.
Information for Nav-O-Matic Autopilots, Electronic Communications and Navigation Equipment are not included in this manual.
These manuals are available from the Cessna Service Parts Center.

iii/(iv blank)

SECTION 1
GENERAL DESCRIPTION

TABLE OF CONTENTS
GENERAL DESCRIPTION ..........
Model 182-Series ............
Description .............

1-1.

GENERAL DESCRIPTION.

1-2.

MODEL 182-SERIES.

Page
1-1
1-1
1-1

1-3. DESCRIPTION. Cessna Model 182-Series aircraft, described in this manual, are high-wing, strutbraced monoplanes of all-metal, semimonocoque construction. These aircraft are equipped with a fixed
tricycle landing gear. Thru aircraft Serial 18260825,
the aircraft employ flat spring-steel main landing
gear struts. Beginning with aircraft Serial 18260826,
the aircraft are equipped with tubular spring-steel
main gear struts. The steerable nose gear is equipped with an air/hydraulic fluid shock strut. Fourplace seating is standard, and a two-place child's
seat may be installed as optional equipment. Model
182-Series aircraft are equipped with a six-cylinder
horizontally opposed, air cooled 0-470-Series Continental engine, driving an all-metal, constant speed
propeller. These aircraft feature rear side windows,
a "wrap around" rear window and a swept-back fin
and rudder.

Aircraft Specifications
Stations ...............
Torque Values ...........

........

1-1
1-1
1-1

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 or computing clearances, remember that such factors as
nose gear strut inflation, tire pressures, tire sizes
and load distribution may result in some dimensions
that are considerably different from those listed.
1-5. STATIONS. A station diagram is shown in
figure 1-2 to assist in locating equipment where a
written description is inadequate or impractical.
1-6. TORQUE VALUES. A chart of recommended
nut torque values is shown in figure 1-3. These
torque values are recommended for all installation
procedures contained in this manual, except where
other values are stipulated. They are not to be
used for checking tightness of installed parts during
service.

1-1

MODELS 182 and A182
GROSS WEIGHT
(Thru 1969 Model 182N)
.................
Take-Off (Thru 1971 Model 182N) .............
Landing (Thru 1971 Model 182N) ...............
(Beginning with 1972 Model 182P) ..............
FUEL CAPACITY
Standard Wing (Total) ..
....
. .
Standard Wing (Usable) .........
Long-Range (Total)
....................
Long-Range (Usable) .................
Standard Wing (Total) . .........
Standard Wing (Usable) ...................
Long-Range (Total) ................
Long-Range (Usable) ..
OIL CAPACITY
(Without External Filter) ..................
(With External Filter)
...................
ENGINE MODEL ........................
PROPELLER (Constant Speed) ................
MAIN WHEEL TIRES (Standard) ..
. ...........
Pressure (Thru 1971 Model 182N) ..............
Pressure (Beginning with 1972 Model 182P) ..
.......
Pressure (Model A182) ...................
MAIN WHEEL TIRES (Optional) .................
Pressure ....................
NOSE WHEEL TIRE (Standard) ..
. ...........
Pressure (Thru 1971 Model 182N) .............
Pressure (Beginning with 1972 Model 182P) . ...
Pressure (Model A182) ................
NOSE WHEEL TIRE (Optional) ..................
Pressure .......................
NOSE GEAR STRUT PRESSURE (Strut Extended)
.........
WHEEL ALIGNMENT
Camber ..
...................
Toe-In . . . . . . . . . . . . . . . . . . . . . .
AILERON TRAVEL
Up .
...............
...
Down ......................
WING FLAP TRAVEL
..
RUDDER TRAVEL (Measured parallel to water line)
Right ..................
....
Left .................
RUDDER TRAVEL (Measured perpendicular to hinge line)
Right ...........................
Left . . . . . . . . . . . . . . . . . . . . . . .
ELEVATOR TRAVEL (Relative to Stabilizer)
Up . . . . . . . . . . . . . . . . . . . . . . . .
Down ...........................
ELEVATOR TRIM TAB TRAVEL
Up . . . . . . . . . . . . . . . . . . . . . . .... .
Down .
.........................
PRINCIPAL DIMENSIONS
Wing Span (Conventional Wing Tip) .............
Wing Span (Conical-Camber Wing Tip) ...........
Tail Span . . . . . . . . . . . . . . . . . . . . ..
Length (Thru 1971 Model 182N) ..............
Length (Beginning with 1972 Model 182P) ..
. .....
Fin Height (Maximum with Nose Gear Depressed and
Flashing Beacon Installed on Fin)
(Thru 1971 Model 182N) ...............
(Beginning with 1972 Model 182P) .......
Track Width (Thru 1971 Model 182N) .............
Track Width (Beginning with 1972 Model 182P) ..
......
BATTERY LOCATION ..
...................
Figure 1-1.
1-2

Change 3

. 2800 lb
. 2950 lb
2800 lb
2950 lb
.

65
60
84
79
.61
56
80
75

.
.

.
.
.
....
.

gal.
gal.
gal.
gal.
gal.1
gal.
gal.
gal.

When not modified by
Cessna Single-Engine
Service Letter SE75-7
and prior to 18262251.
When modified by Cessna
Single-Engine Service
Letter 75-7 and beginning with 18262251.

12 qt
13 qt
CONTINENTAL
82" McCAULEY
6.00 x 6, 6-Ply
32 psi
42 psi
32 psi
8.00 x 6, 6-Ply
25 psi to 35 psi
5. 00 x 5, 6-Ply
50 psi
49 psi
50 psi
6.00 x 6, 4-Ply
30 psi
55 psi to 60 psi

0-470 Series
rating

rating
rating

rating

5° to 7°
. . . . 0" to .06"
20 ° ± 2 °
15 ° ± 2 °
°
. 0O............
to 40 ° , +1° -2 °
24
24

°
°

± 1°
± 1°

27
. . . . 27

°
°

13' ± 1°
13' ± 1°

. . ..
.25
.

26 °
1°
°
17 ± 1°
°
± 2
15 ° ± 1°

36'
35'
11'
. 28'
. 28'

. ...

Aircraft Specifications

2"
10"
8"
1/2"
2" (Add 2" for strobe lights)

. 8' 10-1 1/2"
9' 1-1/2"
7' 11-1/2"
9' 1"
Aft of Baggage Compartment

23.62
I 39.00
71.97

154.00

THRU 182N

0?0

172.00

Figure 1-2. Reference Stations
Change 3

1-3

RECOMMENDED

NUT TORQUES

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

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

TENSION

SHEAR

TORTORQUE

TORQUE

ALT
(NOTE 2)

STD
(NOTE 3)
7-9
12-15
30-40
60-85
95-110
270-300
290-410
480-600
660-780
1300-1500
1500-1800
2200-3300
3000-4200
5400-6600

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

ALT
(NOTE 2)

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

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

(NOTE 5)

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

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

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

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

Figure 1-3.
1-4

Change 2

Torque Values

SECTION 2
GROUND HANDLING, SERVICING, CLEANING, LUBRICATION
TABLE OF CONTENTS

AND INSPECTION

Page

GROUND HANDLING, SERVICING, CLEANING,
LUBRICATION, AND INSPECTION .....
.........
GROUND HANDLING .
Towing ...............
Hoisting ..............
.........
........
.
Jacking
..............
Leveling
Parking ..............
Tie-Down ..............
Flyable Storage ...........
Returning Aircraft to Service . ...
..
Temporary Storage .......
Inspection During Storage .....
Returning Aircraft to Service .....
Indefinite Storage ..........
Inspection During Storage .....
Returning Aircraft to Service .....
SERVICING ..............
........
General Description .
Fuel ................
Carburetor Drain Plug Inspection · · .
Fuel Drains .............
.
.......
Engine Oil ... .
Engine Induction Air Filter ......
.......
Vacuum System Filter .

2-1
2-1
2-1
2-3
2-3
2-3
2-3
2-3
2-3
2-3
2-3
2-4
2-4
2-4
2-5
2-5
2-6
2-6
2-6
2-6
2-6
2-6
2-7
2-7

2-1. GROUND HANDLING, SERVICING, CLEANING,
LUBRICATION AND INSPECTION.

..............
Battery
Tires ...............
Nose Gear Shock Strut ........
Nose Gear Shimmy Dampener .....
Hydraulic Brake System .......
CLEANING ..............
General Description .........
Windshield and Windows . .....
Interior Trim ...........
........
Painted Surfaces
. ..
Aluminum Surfaces .....
........
Engine Compartment
. ....
Upholstery and Interior
..............
Propeller
............
Wheels
LUBRICATION .............
General Description .......
Tachometer Drive Shaft .......
Wheel Bearings ...........
Nose Gear Torque Links .....
Wing Flap Actuator .........
Rod End Bearings . ........
.............
INSPECTION

.
.
.
.
.

2-7
2-7
2-7
2-8
2-8
2-8
2-8
2-8
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-9
2-10
2-17

ground. With the nose wheel clear of the ground, the
aircraft can be turned by pivoting it about the main
wheels.

2-2. GROUND HANDLING.

CAUTION

2-3. TOWING. Moving the aircraft by hand is accomplished by using the wing struts and landing gear
struts as push points. A tow bar attached to the nose
gear should be used for steering and maneuvering the
aircraft on the ground. 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

Figure 2-1.

When towing the aircraft, never turn the nose
wheel more than 30 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.

Tow Bar
Change 3

2-1

ITEM NUMBER

TYPE AND PART NUMBER

REMARKS

Any short jack of capable capacity

Jack

Cessna #SE-767

Universal tail stand (SEE NOTE 1)

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

Universal jack stand (FOR USE WITH ITEM 2)

Cessna #10004-98

Jack point (SEE NOTE 2)

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

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

1. Weighted adjustable stand attaches to tie-down ring.
2. Cessna #10004-98 jack point may be used to raise only one wheel. Do not use brake
casting as a jack point.
3. Items (3). (4), (5) and (6) are available from the Cessna Service Parts Center.
JACKING PROCEDURE
a. Lower aircraft tail so that wing jack can be placed under front spar just outboard of
wing strut.
b. Raise aircraft tail and attach tail stand to tie-down ring. BE SURE that tail stand
weighs enough to keep tail down under all conditions and is strong enough to support
aircraft weight.
c. Raise jacks evenly until desired height is reached.
When using the universal jack point, flexibility of the gear strut will cause the main wheel to slide inboard as the wheel is raised, tilting the jack. The jack must be lowered for a second operation. Jacking
both main wheels simultaneously with universal jack points is not recommended.
Figure 2-2.
2-2

Change 1

Jacking Details

2-4. HOISTING. The aircraft may be lifted with a
hoist of two-ton capacity by using hoisting rings,
which are optional equipment, or by means of suitable slings. The front sling should be hooked to
each upper engine mount at the firewall, and the aft
sling should be positioned around the fuselage at the
first bulkhead forward of the leading edge of the
stabilizer. If the optional hoisting rings are used,
a minimum cable length of 60 inches for each cable
is required to prevent bending of the eyebolt-type
hoisting rings. If desired, a spreader jig may be
fabricated to apply vertical force to the eyebolts.
2-5. JACKING.
procedures.

Refer to figure 2-2 for jacking

2-6. LEVELING. Corresponding points on both
upper door sills may be used to level the aircraft
laterally. Reference point for leveling the aircraft
longitudinally is the top of the tailcone between the
rear window and vertical fin.
2-7. PARKING. Parking precautions depend principally on local conditions. As a general precaution,
set parking brake or chock the wheels and install the
Controls lock. In severe weather and high wind conditions, tie down the aircraft as outlined in paragraph
2-8 if a hangar is not available,
2-8. 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.

Do not set parking brakes during cold weather
when accumulated moisture may freeze the
brakes or when the brakes are overheated.
a. Tie ropes, cables, or chains to the wing tiedown fittings located at the upper end of each wing
strut. Secure the opposite ends of ropes, cables,
or chains to ground anchors.
b. Secure a tie-down rope (no chains or cables) to
upper strut of the nose gear, and secure opposite end
of rope to a 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 external locks may be
installed.
2-9. 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
The aircraft is delivered from Cessna with
a Corrosion Preventive Aircraft Engine Oil
(Military Specification MIL-C-6529 Type II
Rust Ban). This engine oil is a blend of aviation grade straight mineral oil and a corrosion preventive compound. This engine oil
should be used for the first 25 hours of engine
operation. Refer to paragraph 2-21 for oil
changes during the first 50 hours of operation.
During the 30 day non-operational storage or the first
25 hours of intermittent engine operation, the propeller shall be rotated through five revolutions every
seventh day, without running the engine. If the aircraft is stored outside, tie it down in accordance
with paragraph 2-8. 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.
After 30 days, aircraft should be flown for 30 minutes
or ground run-up until oil has reached operating temperature.
2-10. RETURNING AIRCRAFT TO SERVICE. After
flyable storage, returning the aircraft to service is
accomplished by performing a thorough pre-flight
inspection. At the end of the first 25 hours of engine
operation, drain engine oil, clean oil screens and
change external oil filter element. Service engine
with correct grade and quantity of engine oil. Refer

2-11. TEMPORARY STORAGE. Temporary storage
is defined as aircraft in a non-operational status for
a maximum of 90 days. The aircraft is constructed
of corrosion resistant alclad aluminum, which will
last indefinitely under normal conditions if kept clean,
however, these alloys are subject to oxidation. The
first indication of corrosion on unpainted surfaces is
in the form of white deposits or spots. On painted
surfaces, the paint is discolored or blistered. Storage in a dry hangar is essential to good preservation
and should be procured, if possible. Varying conditions will alter the measures of preservation, but
under normal conditions in a dry hangar, and for
storage periods not to exceed 90 days, the following
methods of treatment are suggested.
a. Fill fuel cells with correct grade of gasoline.
b. Clean and wax aircraft thoroughly.
c. Clean any oil or grease from tires and coat tires
with a tire preservative. Cover tires to protect
against grease and oil.
d. Either block up fuselage to relieve pressure on
tires or rotate wheels every 30 days to 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.

Change 2

2-3

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. The
following oils are approved for spraying by
Teledyne Continental Motors: Nude Oil 105Daubert Chemicals Co., 4700 So. Central
Ave., Chicago, Illinois; Petratect VA-Pennsylvania Refining Co., Butler, Pennsylvania,
and Ferro-Gard 1009G-Ranco Laboratories,
Inc., 3617 Brownsville Road, Pittsburgh,
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. If the aircraft is
to be stored outside, stop two-bladed propeller so
that blades are as near horizontal as possible to provide maximum clearance with passing aircraft.
j. Again, spray each cylinder without moving the
crankshaft, to thoroughly cover all interior surfaces
of the cylinder above the piston.
k. Install spark plugs and connect spark plug leads.
1. Apply preservative oil to the engine interior by
spraying approximately two ounces of the preservative
oil through the oil filler tube.
m. Seal all engine openings exposed to the atmosphere, using suitable plugs or non-hygroscopic tape.
Attach a red streamer at each point that a plug or
tape is installed.
n. If the aircraft is to be stored outside, perform
the procedures outlined in paragraph 2-8. 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
2-4

Change 3

effect that the propeller shall not be moved while the
engine is in storage.
2-12. 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 2-11.
2-13. RETURNING AIRCRAFT TO SERVICE. After
temporary storage, use the following procedures 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 value specified
i. Check fuel strainer. Remove and clean filter
screen, if necessary. Check fuel cells and fuel lines
for moisture and sediment. Drain enough fuel to
eliminate moisture and sediment.
j. Perform a thorough pre-flight inspection, then
start and warm-up engine.
2-14. INDEFINITE STORAGE. Indefinite storage
is defined as aircraft in a non-operational status for
an indefinite period of time. Engines treated in accordance with the following may be considered protected against normal atmosphere corrosion, provided the procedures outlined in paragraph 2-15 are
performed at the intervals specified.
a. Operate engine until oil temperature reaches
normal operating range. Drain engine oil sump in
accordance with procedures outlined in paragraph
2-16. Close drain valve or install drain plug.
b. Fill oil sump to normal operating capacity with
corrosion preventive mixture recommended in the
following note. Thoroughly mix and preheat the preventive to a minimum of 221°F at the time it is added
to the engine.

NOTE
Corrosion preventive mixture consists of one
part compound MIL-C-6529C, Type I, mixed
with three parts new lubricating oil of the
grade recommended for service. Continental Motors Corporation recommends Cosmoline No. 1223, supplied by E. F. Houghton
& Co., 305 W. LeHigh Avenue, Philadelphia,
Pa. During all spraying operations, corrosion preventive mixture is preheated to 221 °
to 250°F.
c. Immediately after filling the oil sump with a
corrosion preventive mixture, fly the aircraft for a
period of time not to exceed a maximum of 30 minutes.
d. After flight, with engine operating at 1200 to
1500 rpm, and induction air filter removed, spray
corrosion preventive mixture into induction airbox,
at the rate of one-half gallon per minute. Spray
until heavy black smoke comes from exhaust stack.
Then increase the spray until engine is stopped.
CAUTION
Spraying the mixture too fast can cause a
hydrostatic lock.
e. Do not rotate propeller after completing step
"d. "
f. Remove all spark plugs and spray corrosion
preventive mixture, which has been preheated to
221 ° to 240°F., into all spark plug holes to thoroughly cover interior surfaces of cylinders,
g. Install spark plugs or solid plugs into the lower
spark plug holes and install dehydrator plugs in the
upper spark plug holes. Be sure that dehydrator
plugs are blue in color when installed.
h. Cover spark plug lead terminals with shipping
plugs (AN4060-1), or other suitable covers.
i. With throttle in full open position, place a bag
of desiccant in the induction air intake and seal opening with moisture resistant paper and tape.
j. Place a bag of desiccant in the exhaust tailpipe
and seal openings with moisture resistant tape.
k. Seal cold air inlet to the heater muff with moisture resistant tape.
1. Seal engine breather tube by inserting a protex
plug in the breather hose and clamping in place.
m. Seal all other engine openings exposed to atmosphere, using suitable plugs or non-hygroscopic tape.
NOTE
Attach a red streamer to each location where
plugs or tapes are installed. Either attach
red streamers outside 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 or close drain valve.
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 2-11 thru step "f."
NOTE
As an alternate method of indefinite storage,
the aircraft may be serviced in accordance
with paragraph 2-11, providing the aircraft is
run-up at maximum intervals of 90 days and
then reserviced per paragraph 2-11.
2-15. INSPECTION DURING STORAGE. Aircraft
in indefinite storage shall be inspected as follows:
a. Inspect cylinder protex plugs each 7 days.
b. Change protex plugs if their color indicates an
unsafe condition.
c. If the protex plugs have changed color in one half
of the cylinders, all desiccant material in the engine
should be replaced with new material.
d. Respray the cylinder interiors with corrosion
preventive mixture every 6 months.
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.
2-16. RETURNING AIRCRAFT TO SERVICE. After
indefinite storage, use the following procedure to
return the aircraft to service.
a. Remove aircraft from blocks. Check tires for
correct inflation.
b. Check and install battery.
c. Remove all materials used to seal and cover
openings.
d. Remove warning placards posted at throttle and
propeller.
e. Remove and clean engine oil screen, then reinstall and safety. On aircraft equipped with an external oil filter, install new filter element.
f. Remove oil sump drain plug or open drain valve
and drain sump. Install or close drain valve and
safety.
g. Service and install the induction air filter.
NOTE
The corrosion preventive mixture will mix
with the engine lubricating oil, so flushing
the oil system is not necessary. Draining
the oil sump will remove enough of the corrosion preventive mixture.
h. Remove protex plugs and spark plugs or plugs installed in spark plug holes. Rotate propeller several
revolutions by hand to clear corrosion preventive mixture from cylinders.
i. Clean, gap and install spark plugs. Torque spark
plugs to value specified in Section 11. Connect leads.
j. Check fuel strainer. Remove and clean filter
screen. Check fuel cells and fuel lines for moisture
Change 3

2-5

and sediment. Drain enough fuel to eliminate moisture and sediment.
k. Perform a thorough pre-flight inspection, then
start and warm-up engine.
1. Thoroughly clean and test-fly aircraft.
2-17. SERVICING.
2-18. GENERAL DESCRIPTION. Servicing requirements are shown in figure 2-3. The following paragraphs supplement this figure by adding details not
included in the figure.
2-19. FUEL. Fueltanks should be filled immediately
after flight to lessen condensation in the tanks and lines.
Tank capacities are listed in figure 1-1. The recommended fuel grade to be used is given in figure
2-20. FUEL DRAINS are located at various places
throughout the fuel system. Refer to Section 12 for
location of the various drains in the system. The
strainer drain valve is an integral part of the fuel
strainer assembly. The strainer drain is equipped
with a control which is located adjacent to the oil dipstick. Access to the control is through the oil dipstick access door. Remove drain plugs and open
drain valves at the intervals specified in the inspection charts in this Section. Also, during daily inspection of the fuel strainer, if water is found in the strainer, there is a possibility that the wing tank sumps
or fuel lines contain water. Therefore, all drain
plugs/valves should be removed and all water drained
from the system. To activate drain valve for fuel
sampling, place cup up to valve and depress valve
with rod protruding from cup. (Refer to figure 12-3.)
2-21A. CARBURETOR DRAIN PLUG INSPECTION.
In order to prevent the possibility of thread sealant
contamination in the carburetor float chamber,
cleaning and inspection of the carburetor should be
accomplished at each 100-hour inspection and anytime water in the fuel is suspected.
a. With the fuel valve OFF, remove carburetor
drain plug and clean off any sealant present on the
end of the plug or in the threads on the plug.
b. Inspect drain plug hole in the carburetor and remove any sealant remaining in the hole.
c. Install drain plug as follows:
1. Install drain plug in carburetor 1-1/2 to 2
turns,
2. Apply sealant to drain
Never-Seez RAS-4 or equivalent).
3.
3. Tighten
Tighten and
and safety
safety drain
drain plug.
plug.
f. Turn fuel valve ON and inspect for evidence of
fuel leakage.
2-21. ENGINE OIL. Check engine lubricating oil
2-21. ENGINE OIL. Check engine lubricating oil
with the dipstick five to ten minutes after the engine
has been stopped. The aircraft should be in as near
oil, so that a true reading is obtained.

2-6

Change 3

Engine oil

should be drained while the engine is still hot, and the
nose of the aircraft should be raised slightly for more
positive draining of any sludge which may have collected in the engine oil sump. Engine oil should be
changed every six months, even though less than the
specified hours have accumulated. Reduce these intervals for prolonged operations in dusty areas, in
cold climates where sludging conditions exist, or
where short flights and long idle periods are encountered, which cause sludging conditions. Always
change oil, clean oil screens and clean and/or change
external filter element whenever oil on the dipstick
appears dirty. Aviation grade ashless dispersant oil
conforming to Continental Motors Specification MHS24 and all revisions or supplements thereto and conforming with current Continental Aircraft Engine
Service Bulletins shall be used in the Continental Engines.
NOTE
New or newly overhauled engines should be
operated on aviation grade straight mineral
oil until the first oil change. The aircraft is
delivered from Cessna with straight mineral
oil (MIL-C-6529, Type II, RUST BAN). If
oil must be added during the first 25 hours,
use only aviation grade straight mineral oil
conforming to Specification MIL-6082. After the first 25 hours of operation, drain
engine oil sump and clean both the oil suction
strainer and the oil pressure screen. If an
optional oil filter is installed, change filter
elenent at this time. Refill sump with
straight mineral oil and use until a total of
50 hours have accumulated or oil consumption has stabilized, then change to ashless
dispersant oil.
When changing engine oil, remove and clean oil
screens, or install a new filter element on aircraft
equipped with an external oil filter. An oil quickdrain valve may be installed. This valve provides
a quick and cleaner method of draining the engine oil.
This valve is installed in the oil drain port of the oil
sump. To drain the oil, proceed as follows:
a. Operate engine until oil temperature is at a
normal operating temperature.
b. (With Quick-Drain Valve) Attach a hose to the
quick-drain valve in oil sump. Push upon quickdrain valve until it locks open, and allow oil to
drain through hose into container.
c. (Without Quick-Drain Valve) Remove oil drain
plug from engine sump and allow oil to dran into a
container.
d. After engine oil has drained, close quick-drain
valve, if installed, and remove hose. Install and
safety drain plug.
e. Remove and clean oil screen.
engine oil.

NOTE
Refer to inspection charts for intervals for
changing oil and filter elements. Refer to
figure 2-3 for correct grade of engine oil,
and refer to figure 1-1 for correct capacities.

2-22. ENGINE INDUCTION AIR FILTER. The induction air filter keeps dust and dirt from entering
the induction system. The value of maintaining the
air filter in a good clean condition can never be overstressed. More engine wear is caused through the
use of a dirty or damaged air filter than is generally
believed. The frequency with which the filter should
be removed, inspected, and cleaned will be determined primarily by aircraft operating conditions. A
good general rule however, is to remove, Inspect,
and clean the filter at least every 50 hours of engine
operating time and more frequently if warranted by
operating conditions. Some operators prefer to hold
spare induction air filters at their home base of
operation so that a clean filter is always readily available for use. Under extremely dusty conditions,
daily servicing of the filter is recommended. To
service the induction air 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 cornpressed air a maximum of 30 times or it may
be washed a maximum of 20 times. A new
filter should be installed after using 500 hours
of engine operating time or one year, whichever
should occur first. However, a new filter should
be installed at anytime the existing filter is
damaged. A damaged filter may have sharp
or broken edges in the filtering panels which
would allow unfiltered air to enter the induction system. Any filter that appears doubtful,
shall have a new filter installed in its place.

d. After 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 air box is clean, inspect filter. If
filter is damaged, install a new filter.
f. Install filter at entrance to air box with gasket
on aft face of filter frame and with air flow arrows
on filter frame pointed in the correct direction.
2-23. VACUUM SYSTEM FILTER. The vacuum syster central air filter keeps dust and dirt fron entering
the vacuum operated instruments. Inspect the filter
every 200 hours for damage and cleanliness. Change
central air filter element every 500 hours of operating
time and whenever suction reading drops below 4. 6
inches of mercury. Also, 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.

2-24. BATTERY. Battery servicing involves adding distilled water to maintain the electrolyte even
with the horizontal baffle plate at the bottom of the
filler holes, checking the battery cable connections,
and neutralizing and cleaning off and spilled electrolyte or corrosion. Use bicarbonate of soda (baking
soda) and water to neutralize electrolyte or corrosion. Follow with a thorough flushing with water.
Brighten cables and terminals with a wire brush,
then coat with petroleum jelly before connecting.
The battery box also should be checked and cleaned
if any corrosion is noted. Distilled water, not acid
or "rejuvenators", should be used to maintain electrolyte level. Check the battery every 50 hours (or
at least every 30 days) oftener in hot weather. See
Section 16 for detailed battery removal, installation
and testing.

2-25. TIRES.

Maintain tire pressure at the pressure

specified in figure 1-1. When checking tire pressure,
examine tires for wear, cuts, bruises, and slippage.
Remove oil, grease, and mud from tires with soap
and water.
NOTE
Recommended tire pressures should be maintained. Especially in cold weather, remember
that any drop in temperature of the air inside
a tire causes a corresponding drop in air pressure.

2-26. NOSE GEAR SHOCK STRUT. The nose gear
shock strut requires periodic checking to ensure that
the strut is filled with hydraulic fluid and is inflated
to the correct air pressure. To service the nose gear
shock strut, proceed as follows:

Change 2

2-7

a. Remove valve cap and release air pressure.
b. Remove valve housing.
c. Compress nose gear to its shortest length and
fill strut with hydraulic fluid to the bottom of the
filler hole.
d. Raise nose of aircraft, extend and compress
strut several times to expel any entrapped air, then
lower nose of aircraft and repeat step "c".
e. With strut compressed, install valve housing
assembly.
f. With nose wheel off ground, inflate strut. Shock
strut pressure is listed in figure 1-1.
g. Check strut extension by measuring distance "A",
as indicated in figure 5-5.
NOTE
The nose landing gear shock strut will
normally require only a minimum amount
of service. Maintain the strut extension
pressure as shown in Section 1. Lubricate
landing gear as shown in figure 2-4. 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
lint-free cloth saturated with hydraulic
fluid (MIL-H-5606) or kerosene. All surfaces should be wiped free of excessive
hydraulic fluid.

NOTE
Be sure that the shimmy dampener and
hydraulic fluid are at 70 ° to 80°F while
filling the shimmy dampener.
g. Install filler plug, and wash dampener in cleaning solvent and wipe dry with a clean cloth.
h. Install dampener on 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.
2-28. HYDRAULIC BRAKE SYSTEMS. Check brake
master cylinders and refill with hydraulic fluid as
required every 200 hours. Bleed the brake system
of entrapped air whenever there is a spongy response
to the brake pedals. Refer to Section 5 for filling
and bleeding of the brake systems.
2-29.

2-27. NOSE GEAR SHIMMY DAMPENER. The nose
gear shimmy dampener contains a compensating
mechanism within the hollow piston rod. This is for
thermal expansion and contraction of the hydraulic
fluid in the dampener. The shimmy dampener must
be filled completely with hydraulic fluid, free of entrapped air with the compensating piston bottomed in
the piston rod. Before servicing the shimmy dampener, ascertain that the compensating piston is bottomed in the piston rod. Service the shimmy dampener
at least every 50 hours 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 filler plug to allow excess fluid to escape.
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 of insertion should be 3-13/16
inches.
NOTE
If the wire insertion is less than 3-13/16
inches, the floating piston is lodged in the
shaft. If the wire cannot be used to free
the piston, the rod assembly and piston
should be replaced.
e. After determining that floating piston is bottomed,
move dampener rod to place piston to the end of the
barrel opposite the filler plug.
f. Remove filler plug and fill shimmy dampener
with hydraulic fluid,
2-8

Change 3

CLEANING.

2-30. 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.
2-31. WINDSHIELD AND WINDOWS 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 pressur e. 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.
2-32. INTERIOR TRIM. The instrument panel, interior plastic trim, and control knobs need only be
wiped with a damp cloth. Oil and grease on the control wheels and control knobs can be removed with a
cloth moistened with Stoddard solvent. Volatile solvents, mentioned in the caution note of paragraph 2-31,
must never be used since they soften and craze the
plastic trim.
2-33. PAINTED SURFACES. The painted exterior
surfaces of the aircraft, under normal conditions,
require a minimum of polishing and buffing. Approximately 15 days are required for acrylic or lacquer
paint to cure completely; in most cases, the curing
period will have been completed prior to delivery of
the aircraft. In the event that polishing or buffing is
required within the curing period, it is recommended
that the work be done by an experienced painter.
Generally, the painted surfaces can be kept bright by
washing with water and mild soap, followed by a
rinse with water and drying with cloths or chamois.
Harsh or abrasive soaps or detergents which could
cause corrosion or make scratches should never be
used. Remove stubborn oil and grease with a cloth
moistened with Stoddard solvent. After the curing
period, the aircraft may be waxed with a good automotive wax. A heavier coating of wax on the leading
edges of the wing and tail and on the engine nose cap
will help reduce the abrasion encountered in these
areas.
2-34. ALUMINUM SURFACES. The aluminum surfaces require a minimum of care, but should never
be neglected. The aircraft may be washed with clean
water to remove dirt and may be washed with nonalkaline grease solvents to remove oil and/or grease.
Household type detergent soap powders are effective
cleaners, but should be used cautiously since some
of them are strongly alkaline. Many good aluminum
cleaners, polishes and waxes are available from commercial suppliers of aircraft products,
2-35. ENGINE COMPARTMENT cleaning is essential to minimize any danger of fire, and for proper
inspection of components. The engine compartment
may be washed down with a suitable solvent, such as
Stoddard solvent or equivalent, then dried thoroughly.

longs upholstery fabrics and interior trim. To clean
the interior, proceed as follows:
a. Empty all ash trays and refuse containers.
b. Brush or vacuum clean the upholstery and carpet
to remove dust and dirt.
c. Wipe leather and plastic trim with a damp cloth.
d. Soiled upholstery fabrics and carpet may be
cleaned with a foam-type detergent used according to
the manufacturer's instructions.
e. Oil 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 volatile solvent; it may damage the padding and backing material.
f. Scrape sticky material from fabric with a dull
knife, then spot clean the area.
2-37. 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.
2-38. WHEELS should be washed periodically and
examined for corrosion, chipped paint, and cracks
or dents in the wheel castings. Sand smooth, prime,
and repaint or repair minor defects.
2-39.

2-40. GENERAL DESCRIPTION. Lubrication requirements are shown in figure 2-4. Before adding
lubricant to a fitting, wipe fitting free of dirt. Lubricate until grease appears around part being lubricated, and wipe excess grease from parts. The following paragraphs supplement figure 2-4 by adding details not shown in the figure.
2-41. TACHOMETER DRIVE SHAFT. Refer to Section 15 for details on lubrication of the drive shaft.
2-42. WHEEL BEARINGS. Clean and repack the
wheel bearings at the first 100-hour inspection and
at each 500-hour inspection thereafter. If more
than the usual number of take-offs and landings are
made, extensive taxiing is required, or the aircraft
is operated in dusty areas or under seacoast conditions, cleaning and lubrication of the wheel bearings
shall be accomplished at each 100-hour inspection.
2-43.

Particular care should be given to electrical
equipment before cleaning. Solvent should
not be allowed to enter magnetos, starters,
alternators, voltage regulators, and the like.
Hence, these components should be protected
before saturating the engine with solvent. Any
fuel, oil, and air openings should be covered
before washing the engine with solvent. Caustic
cleaning solutions should be used cautiously and
should always be properly neutralized after
their use.
2-36.

UPHOLSTERY AND INTERIOR cleaning pro-

LUBRICATION.

NOSE GEAR TORQUE LINKS.

Lubricate nose

from a dirt strip or in extremely areas, more frequent lubrication of the torque links is required.
2-44. WING FLAP ACTUATOR.
a. On aircraft prior to Serials 18260698 & A1820136 which have not been modified by Service Kit
SK150-37, proceed as follows:
1. At each 100 hour inspection, inspect wing
flap actuator jack screw and ball retainer assembly
for lubrication, and lubricate if required. Also,
remove, clean and lubricate jack screw whenever
actuator slippage is experienced. If lubrication is
required, proceed as follows:
a. Gain access to actuator by removing
appropriate inspection plates on lower surface of
Change 3

2-9

wing.
b. Expose jack screw by operating flaps to
full-down position.
c. Wipe a small amount of lubricant from
jack screw with a rag and examine for condition.
Lubricant should not be dirty, sticky, gummy or
frothy in appearance.
d. Inspect wiped area on jack screw for
presence of hard scale deposit. Previous wiping
action will have exposed bare metal if no deposit
is present.
e. If any of the preceding conditions exist,
clean and relubricate jack screw as outlined in steps
"f" thru "r".
f. Remove actuator from aircraft in accordance with procedures outlined in Section 7.
g. Remove all existing lubricant from jack
screw and torque tube by running the nut assembly to
the end of the jack screw away from the gearbox, and
soaking the nut assembly and jack screw in Stoddard
solvent.
NOTE
Care must be taken to prevent solvent from
entering gearbox. The gearbox lubricant is
not affected and should not be disturbed.
h. After soaking, clean entire length of jack
screw with a wire brush, rinse with solvent and dry
with compressed air.
NOTE
Do not disassemble nut and ball retainer
assembly.
i. Relubricate jack screw with MIL-G-21164
(Molybdenum Disulfide Grease) as outlined in steps
"j" thru "m".
j. Rotate nut down screw toward the motor.
k. Coat screw and thread end of nut with
grease and run nut to full extension.
1. Repeat the process and pack lubricant in
the cavity between the nut and ball retainer at the

2-10

threaded end of the nut.
m. Repeat the process and work nut back
and forth several times.
n. Remove excess grease.
o. Reinstall actuator in aircraft in accordance with instructions outlined in Section 7.
b. On aircraft prior to Serials 18260698 & A1820136 which have been modified by Service Kit SK15037, proceed as follows:
1. At each 100 hour inspection, expose jack
screw by operating flaps to full-down position, and
inspect wing flap actuator jack screw for proper
lubrication. If lubrication is required, proceed as
follows:
a. Clean jack screw with solvent rag, if
necessary, and dry with compressed air.
b. Relubricate jack screw with MIL-G21164 (Molybdenum Disulfide Grease) as required.
c. On aircraft beginning with Serials 18260698 &
A182-0136, clean and lubricate wing flap actuator
jack screw each 100 hours as follows:
1. Expose jack screw by operating flaps to
full-down position.
2. Clean jack screw threads with solvent rag
and dry with compressed air.
NOTE
It is not necessary to remove actuator from
aircraft to clean or lubricate threads.
3. With oil can apply light coat of No. 10 weight,
non-detergent oil to threads of jack screw.
2-45. ROD END BEARINGS. Periodic inspection
and lubrication is required to prevent corrosion of
the bearing in the rod end. At each 100-hour inspection, disconnect the control rods at the aileron and
nose gear steering bungee, and inspect each rod end
for corrosion. If no corrosion is found, wipe the
surface of the rod end balls with general purpose oil
and rotate ball freely to distribute the oil over its
entire surface and connect the control rods to their
respective units. If corrosion is detected during the
inspection, install new rod ends.

HYDRAULIC FLUID:
SPEC. NO. MIL-H-5606
OXYGEN:
SPEC. NO. MIL-0-27210D
RECOMMENDED FUEL:
ENGINE MODEL 0-470-Series CONTINENTAL
Compliance with conditions stated in Continental aircraft engine Service Bulletins
M74-6 and M75-2 and supplements or revisions thereto, are recommended when
using alternate fuel.
FUEL: 1.
2.
a.
b.

MINIMUM: 80/87 Aviation grade
ALTERNATES:
100/130 Low Lead Avgas (with lead content limited to a maximum of 2 cc Tetraethyl
lead per gallon).
100/130 Higher Lead Avgas (with lead content limited to a maximum of 4. 6 cc Tetraethyl lead per gallon).

Figure 2-3.

Servicing (Sheet 1 of 3)
Change 3

2-11

RECOMMENDED ENGINE OIL:
ENGINE MODEL 0-470-Series CONTINENTAL
AVIATION GRADE:
ABOVE:
BELOW:

SAE 50
SAE 30

Aviation Grade ashless dispersant oil, conforming to Continental Motors Specification MHS-24
and all revisions and supplements thereto, must be used except as noted in paragraph 2-21.
Refer to Continental Aircraft Engine Service Bulletin M75-2 and any superseding bulletins,
revisions or supplements thereto, for further recommendations.

DAILY

FUEL CELLS:
Service after each flight.

Keep full to retard condensation.

Refer to paragraph 2-19.

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

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

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

6
21

Refer to paragraph 2-21 for details.

Check that

PITOT AND STATIC PORTS:
Check for obstructions before first flight of the day.
OXYGEN CYLINDER:
Check for anticipated requirements before each flight.

Refer to Section 14.

(FIRST 25 HOURS

ENGINE OIL SYSTEM:
Refill with straight mineral oil and use until a total of 50 hours have accumulated or oil consumption has stabilized, then change to ashless dispersant oil. Refer to paragraph 2-21.
50 HOURS

INDUCTION AIR FILTER:
Clean filter per paragraph 2-22.

Replace as required.

BATTERY:
Check electrolyte level and clean battery compartment each 50 hours or each 30 days.

ENGINE OIL SYSTEM:
Change oil each 50 hours if engine is NOT equipped with external filter; if equipped with
external filter, change filter element each 50 hours and oil at least at each 100 hours,
or every 6 months.

SHIMMY DAMPENER:
Check fluid level and refill as required in accordance with paragraph 2-27.

Figure 2-3.
2-12

Change 3

Servicing (Sheet 2 of 3)

50 HOURS (Cont.)
TIRES:
Maintain correct tire inflation as listed in figure 1-1.

NOSE GEAR SHOCK STRUT:
Keep strut filled and inflated to correct pressure.

SPARK PLUGS:
Remove, clean and re-gap all spark plugs.

100

Refer to paragraph 2-25.

Refer to paragraph 2-26.

Refer to Section 11.

HOURS

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

CARBURETOR DRAIN PLUG:
Check for thread sealant residue in float chamber.

FUEL STRAINER:
Disassemble and clean strainer bowl and screen.

SELECTOR VALVE DRAIN:
Remove plug and drain off any water or sediment.

Refer to paragraph 2-20.

ALTERNATOR SUPPORT BRACKET:
Check alternator support bracket for security and cracking.
(Also refer to Service Letter SE71-42. )
200

HOURS

VACUUM RELIEF VALVE FILTER:
Change every 100 hours or to coincide with engine overhauls.

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

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

VACUUM SYSTEM CENTRAL AIR FILTER:
Replace every 500 hours.
AS REQUIRED

GROUND SERVICE RECEPTACLE:
Connect to 12-volt DC, negative-ground power unit.

Figure 2-3.

Refer to Section 16.

Servicing (Sheet 3 of 3)
Change 3

2-13

FREQUENCY (HOURS)

METHOD OF APPLICATION

GUN

CAN

(FOR POWDERED
GRAPHITE)

WHERE NO INTERVAL IS SPECIFIED,
LUBRICATE AS REQUIRED AND
WHEN ASSEMBLED OR INSTALLED.
NOTE
The military specifications listed are not mandatory, but are intended as
guides in choosing satisfactory materials. Products of most reputable
manufacturers meet or exceed these specifications.

LUBRICANTS
PG SS-G-659 ............
GR MIL-G-81322A. .........
GH MIL-G-23827A .........
GL MIL-G-21164C. .........
06 MIL-L-7870A ..........
PL VV-P-236 ............
6S MIL-S-8660 ...........
P .................

POWDERED GRAPHITE
GENERAL PURPOSE GREASE
AIRCRAFT AND INSTRUMENT GREASE
HIGH AND LOW TEMPERATURE GREASE
GENERAL PURPOSE OIL
PETROLATUM
DC4 (DOW CORNING)
NO. 10-WEIGHT, NON-DETERGENT OIL

ALSO REFER TO
PARAGRAPH 2-43
NOSE GEAR

ALSO REFER TO

SHIMMY

PARAGRAPH 2-42

DAMPENER
PIVOTS

TORQUE LINKS

WHEEL BEARINGS
WHEEL BEARINGS

Figure 2-4.
2-14

Change 1

Lubrication (Sheet 1 of 3)

OILITE BEARINGS

ELEVATOR DOWN
SPRING LINK RUB
STRIP

CHART IN THIS SECTION AND
TO SECTION 9 OF THIS MANUAL.

RUDDER BARS AND PEDALS

ELEVATOR TRIM
TAB ACTUATOR

BATTERY TERMINALS

PARKING BRAKE
WING STRUT-ATTACH
(UPPER) BOLT & HOLE

HANDLE SHAFT

CABIN DOOR WINDOW
INSERT GROOVES

WING STRUT-ATTACH
(LOWER) BOLT & HOLE*

TRIM WHEEL OILITE AND
NEEDLE BEARINGS

* UPON INSTALLATION
Figure 2-4.

Lubrication (Sheet 2 of 3)
Change 1

2-15

CONTROL COLUMN

NEEDLE
BEARINGS

NEEDLE BEARING

THRUST BEARING

NEEDLE BEARING

ELECTRIC FLAP

NEEDLE BEARINGS
ROD END
BEARINGS

NOTES
Sealed bearings require no lubrication.
Do not lubricate roller chains or cables except under seacoast conditions.
dry cloth.

Wipe with a clean,

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

Figure 2-4.

Lubrication (Sheet 3 of 3)

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.
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

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.

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.

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.

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.

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

Change 3

2-17

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 m 1. (a) and (b).
Cessna Progressive Care may be utilized as a total concept program which
insures that the inspection intervals in the inspection charts are not exceeded.
Manuals and forms which are required for conducting Progressive Care inspections are available from the Cessna Service Parts Center.

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.
To be made available upon request:
1. Aircraft Log Book and Engine Log Book.

2-18

Change 1

(h)

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

Engine temperatures and pressures.
Static rpm. (Also refer to Section 11 of this Manual).
Magneto drop. (Also refer to Section 11 of this Manual).
Engine response to changes in power.
Any unusual engine noises.
Fuel selector and/or shut-off valve; operate engine(s) on each tank (or cell) position
and OFF position long enough to ensure shut-off and/or selector valve functions
properly.
Idling speed and mixture; proper idle cut-off.
Alternator and ammeter.
Suction gage.
Fuel flow indicator.

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

SHOP NOTES:

Change 1

2-19

SPECIAL INSPECTION ITEM
EACH 200 HOURS

IMPORTANT

EACH 100 HOURS

READ ALL INSPECTION REQUIREMENTS PARAGRAPHS PRIOR TO
USING THESE CHARTS.

EACH 50 HOURS

PROPELLER
1.

Spinner

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

Blades.

Bolts and nuts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Hub ........................

Governor and control . . . . ...

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

. . . .

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

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

. . . . . .....

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

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

Oil cooler ...............................

Induction air filter .........................................

Induction airbox, air valves, doors and controls .......

Cold and hot air hoses . ...............

Engine baffles

Cylinders, rocker box covers and push rod housings

Crankcase, oil sump, accessory section and front crankshaft seal ..........

Hoses, metal lines and fittings

.....

1
.

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

.....

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

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, oil separator and relief valve

17.

Vacuum relief valve filter ........................

18.

Engine controls and linkage

19.

Engine shock mounts, mount structure and ground straps ........

2-20

Change 1

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

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

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

...........

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

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

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
SPECIAL INSPECTION ITEM
EACH 200 HOURS
EACH 100 HOURS
EACH 50 HOURS
20. Cabin heat valves, doors and controls ....................................................................................
21. Starter, solenoid and electrical connections...........................................................................
22. Starter brushes, brush leads, commutator.............................................................................
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.......................................................................................................................
28. Carburetor and drain plug (Refer to Service Letter (SE73-13.) .............................................

29. Firew all...................................................................................................................................
30. Engine cowl flaps and controls...............................................................................................
31. E ng ine cow ling .......................................................................................................................
32. Cowl flap hinges and hinge pins (Refer to Service Letter SE71-27.)..................

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

33. Carburetor throttle arm attachment (Refer to Service Letter SE71-17.) ................................
34. Alternator support bracket for security (Refer to Service Letter SE71-42.)............................
FUEL SYSTEM
1. Fuel strainer, drain valve and control, fuel cell vents, caps and placards..............................
2. Fuel strainer screen and bowl ................................................................................................
3. Drain fuel and check cell interior, attachment and outlet screens..........................................
4. Fuel cells and sump drains.....................................................................................................

5. Fuel selector valve and placards (Refer to Service Letter SE74-1.) ........................ ...........
6. E ng ine prim er .........................................................................................................................
7. Fuel quantity indicators and transmitters................................................................................
8. Perform a fuel quantity indicating system operational test. Refer to
Section 15 for detailed accomplishment instructions ............................................................

LANDING GEAR
1. Brake fluid, lines and hose, linings, discs, brake assemblies and master cylinders ...............
2. Main gear wheels ........................................
.....................................
3. W heel bearings........................................................................................................................

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

Revision 4
Mar 1/2004

2-21
© Cessna Aircraft Company

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

Tires

Torque link lubrication ......................

Parking brake system

Nose gear strut and shimmy dampener (service as required)

Nose gear wheel . ...

. . . . . .

. . . . ..

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

. . .

...

. ...

. . . . . .

....

. .

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

10.

Nose gear fork

11.

Nose gear steering system

12.

Park brake and toe brakes operational test

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

AIRFRAME

. . . . . .

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

Aircraft exterior

Aircraft structure

Windows, windshield, doors and seals

Seat stops, seat rails, upholstery, structure and mounting .

Seat belts and shoulder harnesses

Control column bearings, pulleys, cables and turnbuckles

Control lock, control wheel and control column mechanism . ....

Instruments and markings ....................

Gyros central air filter .....................

(Refer to Service Letters SE72-3 and SE72-29.) .

.......

.
. . . . . . .

. . .........

0
. . . .

10
5

Magnetic compass compensation .................

11.

Instrument wiring and plumbing

12.

Instrument panel, shock mounts, ground straps, cover, decals and la beling

13.

Defrosting, heating and ventilating systems and controls ......

14.

Cabin upholstery, trim, sunvisors and ash trays..........

15.

Area beneath floor, lines, hose, wires and control cables

16.

Lights, switches, circuit breakers, fuses and spare fuses

.....

17.

Exterior lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

18.

Pitot and static systems

19.

Stall warning unit and pitot heater ...

20.

Radios, radio controls, avionics and flight instruments
Change 3

. . .

10.

2-22

.......

.....

. . . . . . . .

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

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

......

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

21.

Antennas and cables

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

22.

Battery, battery box and battery cables ..

23.

Battery electrolyte

24.

Emergency locator transmitter

25.

Oxygen system

26.

Oxygen supply, masks and hose

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

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

.........

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

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

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

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 .

Chains, terminals, sprockets and chain guards

Trim control wheels, indicators, actuator and bungee

Travel stops

Decals and labeling ...

Flap control switch, flap rollers and flap position indicator .............

Flap motor, transmission, limit switches, structure, linkage, bellcranks, etc.

Flap actuator jackscrew threads .........................

Elevators, trim tab, hinges and push-pull tube

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

.
.

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

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

14
.................

10.

Elevator trim tab actuator lubrication and tab free-play inspection

11.

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

12.

External skins of control surfaces and tabs

13.

Internal structure of control surfaces

14.

Balance weight attachment

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

.........

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

Change 1

2-23

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
SPECIAL INSPECTION ITEMS
1.

First 25 hours; (refill with straight mineral oil and use until a total of 50 hours has accumulated or oil
consumption has stabilized) then change to ashless dispersant oil. Change oil each 50 hours if engine is
NOT equipped with external oil filter; if equipped with an external oil filter, change filter element at each
50 hours and oil at each 100 hours or every six months.

Clean filter per paragraph 2-22. Replace as required.

Replace engine compartment hoses (Cessna-installed only) every 5 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 replaced within 120 days after
receiving the new hose(s) from Cessna. Replace drain hoses on condition. Engine flexible hoses,
(Continental Motors installed) refer to Continental Motors Maintenance Manual and Continental Motors
Engine Service Bulletins.

General inspection every 50 hours. Refer to Section 11 for 100 hour inspection.

Each 1000 hours, or to coincide with engine overhaul.

Each 50 hours for general condition and freedom of movement. These controls are not repairable.
Replace as required at each engine overhaul.

Each 500 hours.

Internal timing and magneto-to-engine timing limits are described in detail in Section 11.

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

10. Replace each 500 hours.
11. Check electrolyte level and clean battery compartment each 50 hours or 30 days.
12. Refer to Section 16 of this Service Manual.
13. Inspect masks, hose fittings for condition, routing and support. Test, operate and check for leaks.
14. Refer to paragraph 2-44 for detailed instructions for various serial ranges.
15. Lubrication of the actuator is required each 1000 hours and/or 3 years, whichever comes first. Refer to
Figure 2-4 for grease specification.
NOTE:

Refer to Section 9 of this manual for free-play limits, inspection, replacement and/or repair.

16. Fuel quantity indicating system operational test is required every 12 months. Refer to Section 15 for
detailed accomplishment instructions.
NOTE:

A high time inspection is merely a 100-hour inspection with the addition of an engine overhaul.
Refer to Teledyne Continental Service Information Letter SIL98-9, or latest revision, for
recommended time between overhaul for 0-470 series engines. At the time of overhaul, engine
accessories should be overhauled.
Propeller overhaul should coincide with engine overhaul, but intervals between overhauls of the
propeller shall not exceed 1200 hours, except as stipulated in current issues of the McCauley
Accessory Division Service Information Summary and currently effective Service Manuals,
Bulletins and Letters.

2-24

Revision 4
©Cessna Aircraft Company

Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
2-46.

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 CFR 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 CFR 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 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
COMPONENT

REPLACEMENT
TIME

OVERHAUL

Restraint Assembly Pilot Copilot
And Passenger Seats

10 years

Trim Tab Actuator

1,000 hours or 3 years,
Whichever occurs first

YES

Vacuum System Filter

500 hours

Vacuum System Hoses

10 years

Pitot and Static System Hoses

10 years

Vacuum Relief/Regulator Valve
Filter (If Installed)

500 hours

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

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

Revision 4
Mar 1/2004

2-25
© Cessna Aircraft Company

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
COMPONENT

REPLACEMENT
TIME

OVERHAUL

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

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

Engine Air Filter

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

Engine Mixture, Throttle, and
Propeller Controls

At engine TBO

Oxygen Bottle - Lightweight Steel

Every 24 years or 4380
cycles, whichever occurs first

(ICC-3HT, DOT-3HT)
Oxygen Bottle - Composite

Every 15 years

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

6 Years or at vacuum
pump replacement,
whichever occurs first

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

500 hours
(Note 10)

Standby Dry Vacuum Pump

500 hours or 10 Years,
whichever occurs first
(Note 10)

(DOT-E8162)

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

REPLACEMENT
TIME

OVERHAUL

ELT Battery

(Note 3)

Vacuum Manifold

(Note 4)

Magnetos

(Note 5)

YES

Engine

(Note 6)

YES

Engine Flexible Hoses
(TCM-Installed)

(Note 2)

Auxiliary Electric Fuel Pump

(Note 7)

YES

Propeller

(Note 8)

YES

2-26
© Cessna Aircraft Company

Revision 4
Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE 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 fluidcarrying 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:

For TCM engines, 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 over service the air filter. Over servicing 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 comes first.
For a vacuum pump equipped with a wear indicator, replace pump according to the vacuum pump
manufacturer's recommended inspection and replacement intervals.

Revision 4
Mar 1/2004

2-27/ (2-28 Blank)
© Cessna Aircraft Company

SECTION 3
FUSELAGE

Page

TABLE OF CONTENTS
FUSELAGE ................
Windshield and Windows
..
....
Description .............
Cleaning .............
Waxing ..............
Repairs ..............
Scratches .............
Cracks .............
Windshield .............
Removal
..
. .....
Installation
..........
Windows
..............
Movable .
...........
Removal and Installation
Wrap-Around Rear .......
Removal and Installation
.........
Overhead ..
Removal and Installation
Fixed
.........
...
Cabin Doors ..............
Removal and Installation .....
Adjustment ............
Weatherstrip. ...........
Wedge Adjustment .........
Latches ..............
Description ..........
Adjustment ..........
Lock ..............
Indexing Inside Handle ......
Baggage Door .............
Removal and Installation .....
. ..
Assist Straps ............
3-1.

3-1
3-1
3-1
3-1
3-1
3-1
.
3-1
.
3-2
3-4
.
3-4
3-4
3-4
3-4
. .3-4
3-4
3-4
.
3-4
. .3-4
.
3-4
3-4
3-4
.
3-4
3-4
3-4
3-7
3-7
3-7
3-7
.
3-7
3-7
.
3-7
. 3-7
.

FUSELAGE.

3-2. WINDSHIELD AND WINDOWS.
3-3. DESCRIPTION. The windshield and windows
are single-piece acrylic plastic panels set in sealing
strips and held by formed retaining strips secured
to the fuselage with screws and rivets. Presstite No.
579. 6 sealing compound used in conjunction with a
felt seal is applied to all edges of windshield and
windows with exception of 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
further abrasion. Use a good grade of commercial
wax applied in a thin, even coat. Bring wax to a
high polish by rubbing lightly with a clean, dry flannel cloth.

Removal and Installation ..
..
. 3-71
Seats ..
. ...........
.
3-7
.......
3-7
Pilot and Copilot ..
Reclining Back .........
3-7
Vertical Adjust/Reclining Back .3-7
Articulating Recline/Vertical
Adjust .
...........
3-7
Description ..
......
3-7
Removal and Installation ..
3-7
3-7
Center ..
. .........
Double-Width Bottom/Individual
Reclining Backs .......
3-7
Description .
.......
3-7
Removal and Installation ..
3-7
Auxiliary .............
3-20
. 3-20
Fold-Up ...........
3-20
Description ........
Removal and Installation . . 3-20
3-20
Repair .............
3-20
Cabin Upholstery ............
3-20
Materials and Tools ..........
Soundproofing .............
3-20
3-20
Cabin Headliner ............
. 3-21
Removal and Installation .....
. 3-21
Upholstery Side Panels ........
Carpeting ...............
3-21
Safety Provisions
.........
. 3-21
Cargo Tie-Downs .........
3-21
Safety Belts ...........
.
3-21
Shoulder Harness
...........
3-22
. .3-22
Glider Tow Hook ...........
3-22
Rear View Mirror ...........
3-6. REPAIRS. Damaged window panels and windshield may be removed and replaced if damage is
extensive. However, certain repairs as prescribed
in the following paragraphs can be made successfully
without removing damaged part from aircraft. Three
types of temporary repairs for cracked plastic are
possible. No repairs of any kind are recommended
on highly-stressed or compound curves where repair
would be likely to affect pilot's field of vision.
Curved areas are more difficult to repair than flat
areas and any repaired area is both structurally and
optically inferior to the original surface.
3-7. SCRATCHES. Scratches on clear plastic surfaces can be removed by hand-sanding operations
followed by buffing and polishing, if steps below are
followed carefully.
a. Wrap a piece of No. 320 (or finer) sandpaper or
abrasive cloth around a rubber pad or wood block.
Rub surface around scratch with a circular motion,
keeping abrasive constantly wet with clean water to
prevent scratching surface further. Use minimum

Change 2

3-1

WOOD REINFORCEMENT

WOOD

ALWAYS DRILL END OF CRACK CUSHION OF
TO RELIEVE STRAIN
OR
OR FABRIC
FABRICC
WRONG

RIGHT

SOFT WIRE
LACING

CEMENTED
FABRIC PATCH
TEMPORARY
OF CRACKS

Figure 3-1.

Repair of Windshield and Windows

pressure and cover an area large enough to prevent
formation of "bull's-eyes" or other optical distortions.
CAUTION
Do not use a coarse grade of abrasive.
320 is of maximum coarseness.

No.

b. Continue sanding operation, using progressively
finer grade abrasives until scratches disappear.
c. When scratches have been removed, wash area
thoroughly with clean water to remove all gritty particles. The entire sanded area will be clouded with
minute scratches which must be removed to restore
transparency.
d. Apply fresh tallow or buffing compound to a
motor-driven buffing wheel. Hold wheel against plastic surface, moving it constantly over damaged area
until cloudy appearance disappears. A 2000-foot-perminute surface speed is recommended to prevent
overheating and distortion. (Example: 750 rpm
polishing machine with a 10 inch buffing bonnet.)
NOTE
Polishing can be accomplished by hand but
will require a considerably longer period
of time to attain the same result as produced by a buffing wheel.
e. When buffing is finished, wash area thoroughly
and dry with a soft flannel cloth. Allow surface to
cool and inspect area to determine if full transparency has been restored. Apply a thin coat of hard
wax and polish surface lightly with a clean flannel
cloth.
3-2

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

Detail B

Detail C
3

to all edges of windshield and windows when
felt sealing strip (3) is used.
TYPICAL METHODS OF RETAINING FIXED WINDOWS

Figure 3-2.

Windshield and Fixed Window Installation
3-3

lacing material. This type of repair is used as a
temporary measure ONLY, and as soon as facilities
are available, panel should be replaced.
WINDSHIELD.

3-9.

(Refer to figure 3-2.)

3-10. REMOVAL.
a. Drill out rivets securing front retainer strip.
b. Remove wing fairings over windshield edges.
NOTE
Remove and tape compass clear of work
area. Do not disconnect electrical wiring.
c. Pull windshield straight forward, out of side
and top retainers. Remove top retainer if necessary.
3-11. 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 installation.
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.
If not previously removed, top retainer may be removed if necessary. Starting at upper corner and
gradually working windshield into position is recommended.
NOTE
Screws and self-locking nuts may be used
instead of rivets which fasten front retaining
strip to cowl deck. If at least No. 6 screws
are used, no loss of strength will result.
3-12.

WINDOWS.

3-13. MOVABLE. (Refer to figure 3-3.) A movable
window, hinged at the top, is installed in the left cabin door thru 1975 models and beginning with 1976 may
also be installed in the RH door. Beginning with 1974
models a close fitting window frame is employed with
an improved seal. The seal is attached to the door
frame using EC-880 (3-M Company) or equivalent.
3-14. 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.
3-15. WRAP-AROUND REAR. The rear window is
a one-piece acrylic plastic panel set in sealing strips
and held in place by retaining strips.
3-16. REMOVAL AND INSTALLATION.
a. Remove upholstery as necessary to expose retainer strips inside cabin.
b. Drill out rivets as necessary to remove retainers on both sides and lower edge of window.
3-4

Change 3

c. Remove window by starting at aft edge and
pulling window into 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, use care not to crack panel and
file or grind away excess plastic.
3-17. OVERHEAD. (Refer to figure 3-2.) Overhead
cabin windows, located in the cabin top, may be installed. These windows are one-piece acrylic plastic
panels set in sealing strips and held in place by retaining strips.
3-18. REMOVAL AND INSTALLATION.
a. Remove headliner and trim panels.
b. Drill out rivets as necessary to remove retainer
strips.
c. Reverse preceding steps for reinstallation. Apply
felt strip and sealing compound to all edges of window
to prevent leaks. Check fit and carefully file or grind
away excess plastic. Use care not to crack plastic
when installing.
3-19. FIXED. (Refer to figure 3-2.) Fixed windows, mounted in sealing strips and sealing compound, are held in place by various retainer strips.
To replace side windows, remove upholstery and
trim panels as necessary and drill out rivets securing retainers. Apply felt strip and sealing compound
to all edges of window to prevent leaks. Check fit
and file or grind away excess plastic. Use care not
to crack plastic when installing.
3-20.

CABIN DOORS.

(Refer to figure 3-3.)

3-21. REMOVAL AND INSTALLATION. Removal
of cabin doors is accomplished either by removing
screws which attach hinges and door stop or by
removing hinge pins attaching hinges and door stop.
If permanent hinge pins are removed, they may be
replaced by clevis pins secured with cotter pins
or new hinge pins may be installed and "spin-bradded. " When fitting a new door, some trimming of
door skin at edges and some reforming with a soft
mallet may be necessary to achieve a good fit.
3-22. ADJUSTMENT. Cabin doors should be adjusted so skin fairs with fuselage skin. Slots at the
latch plate permit re-positioning of striker plate.
Depth of latch engagement may be changed by adding
or removing washers or shims between striker plate
and doorpost.
3-23. WEATHERSTRIP. Rubber extruded seals are
installed around the edges of the door. Beginning
with serial 18263830 an improved type door seal is
used which has a hollow center and small flutes extending along its length. When replacing door seals
ensure mating surfaces are clean, dry and free of
oil and grease. Position butt ends of seal at door
low point and cut a small notch in the hollow seal
for drainage. Apply a thin, even coat of EC-880
adhesive (3-M Co. ) or equivalent to each surface
and allow to dry until tacky before pressing into place.
3-24.

WEDGE ADJUSTMENT.

Wedges at upper for-

NOTE
8

L-

on the bonded door, as forming of the
flanges could cause damage to the bond-

REFER TO FIGURE 3-4

a:1

DetailD

Detail

with Aircraft Serial 18261426 and
Detail C
1. Upholstery Clip
2. Upholstery Panel
3. Wedge
4. Spring
5. Window Stop
6. Window Hinge

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

Nut
Lock Assembly
Latch Assembly
Door Stop Arm
Spring-Loaded Plunger
Wedge

23.
24.
25.
26.
27.
28.

7. Latch Plate

18.

Spacer

29.

Clamp

19.
20.
21.
22.

Stop Assembly
Reinforcement
Hinge
Pin

30.

Window Moulding

8. Cabin Door
9. Window Frame
10. Window
11. Washer

Lower Hinge
Upper Hinge
Door Jamp
Screw
Pull Handle
Clamp Cover

THRU AIRCRAFT SERIAL 18262465

Figure 3-3.

Cabin Door Installation (Sheet 1 of 2)
Change 2

3-5

NOTE

Beginning with serial 18264296
an openable window may be installed in the RH cabin door.
Procedures are similar to door
illustrated.

3-4
NOTE

since damage may occur to the

silicone grease.
Spray cabin door and window
seals with MS-122 (18598) or
DOOR INSTALLATION
equivalent. Caution, do not
BEGINNING WITH AIRCRAFT
overspray, confine to the seal.
SERIAL 18262466

NOTE
Trim cutout in inner door pan if
necessary to maintain . 10 minimum
clearance with door stop arm.

Figure 3-3. Cabin Door Installation (Sheet 2 of 2)
3-6

Change 3

AIRCRAFT SERIALS 18260111 THRU
18262465 AND A182-0117 AND ON
**

13
14

AIRCRAFT SERIALS 18259900 THRU\
18262465 AND A182-0117 AND ON

* THRU AIRCRAFT SERIALS 18259899 AND
A182 -0116
* *THRU AIRCRAFT SERIAL 18261425 AND
BEGINNING WITH A182-0117

*BEGINNING

WITH AIRCRAFT SERIAL
18261426 THRU 18262465

6 7

ROTATED 90 °

11.

2.
....

13.

4.
5.
6.
7.

Top Bolt Guide
Bolt
Side Bolt Guide
Base Bolt Guide
Latch Base Plate
Abrasive Pad
Lockplate

8. Bracket

22
P

25
J

26
27
NOTE
otary clutch components
are matched upon assembly.
The clutch mechanism, if
defective, should be replaced
as a unit.

CABIN DOOR
ROTARY CLUTCH

THRU AIRCRAFT SERIAL 18262465

Figure 3-4.

9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
2
23.
24
25.
26.
27.
28.
29.

Spring
Nylon Washer
Placard
Escutcheon
Placard
Inside Handle
Clip
Plate Assembly
Support
Shaft Assembly
Bolt Push Rod
Outside Handle
Pull Bar
Mounting Structure
Shim
Rotary Clutch
Guide
Door Post
Cover
Handle Adjust Screw
Bolt Adjust Screw

Door Latch Installation (Sheet 1 of 2)
Change 1

3-6A

NOTE
3-27 for bolt (item 2)

SERIAL 18264296
2

Set adjustment screw (29)
in the slot to maintain door
handle 8 15' above center
door
locked position.27
is
in

Figure 3-4.
3-6B

Change 3

r-1
BEGINNING WITH AIRCRAFT

Rotary clutch components are
matched upon assembly. The
clutch mechanism, if defec-

the

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

3-25.

LATCHES.

(Refer to figure 3-4.)

3-26. DESCRIPTION. The cabin door latch is a
push-pull bolt type, utilizing a rotary clutch for
positive bolt engagement. As door is closed, teeth
on underside of bolt engage gear teeth on clutch.
The clutch gear rotates in one direction only and
holds door until handle is moved to LOCK position,
driving bolt into slot.
NOTE
On some aircraft the bolt will have a notch
in the aft end to allow for a better contour
fit between door and fuselage.
3-27. ADJUSTMENT. Vertical adjustment of the rotary clutch is afforded by slotted holes which ensures sufficient gear-to-bolt engagement and proper
alignment. The extension or retraction of the bolt item
(2) is controlled by adjusting mounting bolts item (29)
in the slotted holes. Lossen screws sufficient to move
latch base forward on the door to retract bolt and aft
Close door carefully after adjustment and
check clearance between bolt and door jamb
and clutch engagement.
3-28. LOCK. In addition to interior locks, a cylinder and key type lock is installed on left door. If
lock is to be replaced, the new one may be modified
to accept the original key. This is desirable, as the
same key is used for ignition switch and cabin door
lock. After removing old lock from door, proceed
as follows:
a. Remove lock cylinder from new housing.
b. Insert original key into new cylinder and file off
any protruding tumblers flush with cylinder. Without
removing key, check that cylinder rotates freely in
housing.
c. Install lock assembly in door and check lock
operation with door open.
d. Destroy new key and disregard code number on
cylinder.
3-29. INDEXING INSIDE HANDLE. (Refer to figure
3-4.) When inside door handle is removed, reinstall
in relation to position of bolt (2) which is springloaded to CLOSE position. The following procedure
may be used:
a. Temporarily install handle (14) on shaft assembly (18) approximately vertical.
b. Move handle (14) back and forth until handle centers in spring-loaded position.
c. Without rotating shaft assembly (18), remove
handle and install placard (11) with CLOSE index at
top and press placard to seat prongs.
d. Install nylon washer (10).
e. Install handle (14) to align with CLOSE index on
placard (11) and install clip (15).
f. Ensure bolt (2) clears doorpost and teeth engage
clutch gear when handle (14) is in CLOSE position.
Beginning with 1974 models the inside handle is moved
forward on the door and fits into the armrest when it
is moved to the locked position. Install the handle on
the serated shaft so that the forward end of the handle

is 8° 15' above the centerline of the handle shaft when
in the locked position. A small amount of adjustment
can be accomplished by loosening the shaft mounting
bolts and moving bolt item (28) in the slot to raise or
lower the forward end of the handle.
3-30.

BAGGAGE DOOR.

(Refer to figure 3-5.)

3-31. REMOVAL AND INSTALLATION.
a. Disconnect door-stop chain (9).
b. Remove inside door handle (2) if installed.
c. Remove screws securing upholstery panel and
remove panel.
d. Remove bolts (11) securing door to hinges or
remove clevis pins (10) securing hinges to brackets.
e. Reverse preceding steps for reinstallation.
3-31A. ASSIST STRAPS (Refer to figure 3-3)
3-31B. REMOVAL AND INSTALLATION. Figure
3-3 may be used as a guide for removal and installtion of the assist straps.

PILOT AND COPILOT.
a. RECLINING BACK.
b. VERTICAL ADJUST/RECLINING BACK.
c. ARTICULATING RECLINE/VERTICAL
ADJUST.
3-33.

3-34. DESCRIPTION. These seats are manuallyoperated throughout their full range of operation.
Seat stops are provided to limit fore-and-aft travel.
3-35. REMOVAL AND INSTALLATION.
a. Remove seat stops from rails.
b. Slide seat fore-and-aft to disengage seat rollers
from rails.
c. Lift seat out.
d. Reverse preceding steps for reinstallation. Ensure all seat stops are reinstalled.

WARNING
It is extremely important that pilot's seat
stops are installed, since acceleration and
deceleration could possible permit seat to
become disengaged from seat rails and
create a hazardous situation, especially during take-off and landing.
CENTER.
a. DOUBLE-WIDTH BOTTOM/INDIVIDUAL
RECLINING BACKS.
3-36.

3-37. DESCRIPTION. These seats are permanently
bolted to the cabin structure and incorporate no adjustment provisions other than manually-adjustable
three position backs.
3-38. REMOVAL AND INSTALLATION.
a. Remove bolts securing seat to cabin structure.
b. Lift seat out.
c. Reverse preceding steps for reinstallation.

Change 2

3-7

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

"""

J/j

Detail A

10
NOTE
AIRCRAFT SERIALS 18260446 AND ON
AND A182-0138 AND ON INCORPORATE
A BONDED BAGGAGE DOOR.

* Use spacer and shims (6)
as required to align outside handle (7) flush with
door skin.

~year,

* Beginning with 1971 Model
inside handle (2) is
not installed.

Detail B
Figure 3-5.
3-8

Baggage Door Installation

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

Baggage Door
Inside Handle
Cam
Latch Assembly
Lock Assembly
Shim or Spacer
Outside Handle
Striker Plate
Chain
Clevis Pin
Bolt
Hinge

PILOT AND COPILOT SEAT
(STANDARD THRU 1972)

RECLINING BACK

Figure 3-6.

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

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

Seat Installation (Sheet 1 of 8)
Change 1

3-9

PILOT AND COPILOT SEAT
(STANDARD BEGINNING WITH 1973)

BEGINNING WITH SERIAL 18264296
12

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

Figure 3-6.
3-10.

Change 3

Seat Installation (2 of 8)

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

PILOT AND COPILOT SEAT
(OPTIONAL 1969)

RECLINING BACK

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

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

Figure 3-6.

Seat Installation (Sheet 3 of 8)
Change 1

3-11

PILOT AND COPILOT SEAT
(OPTIONAL 1970 THRU 1972)
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.

Vertical Adjustment Handle
Adjustment Pin
Fore/Aft Adjustment Handle
Seat Bottom
Articulating Adjustment Handle
Bellcrank
Adjustment Screw
Seat Back
Trim Bracket
Spacer
Channel
Torque Tube

14.

Roller

* NOTE
The nut on adjustment screw (7)
is rotated 180 beginning with
SERIAL 18260826 AND ON.

Detail A

Detail

B
Figure 3-6.

3-12

Change 1

Seat Installation (Sheet 4 of 8)

PILOT AND COPILOT SEAT
OPTIONAL THRU
1973 MODELS

ARTICULATING RECLINE/
VERTICAL ADJUST

Detail

Detail B

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

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

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

Figure 3-6.

Bellcrank
Seat Back
Spacer
Channel
Torque Tube
Seat Structure
Roller

Seat Installation (Sheet 5 of 8)
Change 1

3-13

PILOT AND COPILOT SEAT
BEGINNING WITH 1974 MODELS
(OPTIONAL INSTALLATION)
9

ARTICULATING BACK/
VERTICAL ADJUST

Detail A

SERIALS 18263080 THRU 18264295
BEGINNING WITH SERIALS 18264296

Detail B

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

Vertical Adjustment Handle
Fore/Aft Adjustment Handle
Adjustment Pin
Spring
Seat Bottom
Articulating Adjustment Handle
Adjustment Screw
Bellcrank

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

Figure 3-6.
3-14

Change 3

Seat Back
Spacer
Channel
Torque Tube
Seat Structure
Roller
Stiffner
Seat Belt Retainer

Seat Installation (Sheet 6 of 8)

CENTER SEAT
(STANDARD)
1.
2.
3.
4.
5.
6.
7.
8.

Seat Bottom
Spring
Bushing
Seat Back
Recline Cam
Recline Handle
Recline Pawl
Control Shaft

DOUBLE-WIDTH BOTTOM/

Figure 3-6.

Seat Installation (Sheet 7 of 8)
Change 1

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

AUXILARY SEATS
(OPTIONAL)
THRU AIRCRAFT SERIALS
18260055 AND A182-0116

Detail 7 A
BEGINNING WITH
18260446

3.
4.
5.

Strap
Hinge Bracket
Seat Bottom Structure

Change 1

3-15

SEAT BACK (REF)

CLEVIS BOLT (REF)

2.50" R. (CONSTANT AT EACH NOTCH)

1414230-2 (FULL
1414111-5 (VERTICALLY
ADJUSTABLE SEAT)

REPLACEMENT PROCEDURE:
a.

Remove seat from aircraft.

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

Drill out existing rivets and insert new cam assembly (2). Position seat back so pawl (3) engages first cam slot as illustrated.

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

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

Reinstall upholstery, upholstery panels and seat.

Figure 3-7.
3-16

Seat Back Cam Replacement

18259306 THRU 18260445

Detail A

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

18262679
Detail

Screw
Hook
Hook
Shoulder Harness
Spacer
Washer

6.
7.
7.
8.
9.
10.

Cover
Bolt
Bolt
Clip
Eye Bolt
Nut

Figure 3-8.

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

Bracket
Seat
Frame
Seat Frame
Seat Belt
Spacer
Harness Tray

16.
17.
17.
18.
19.

Inertia
Reel
Plate Trim
Plate Trim
Attaching Plate
Inertia Reel Cover

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

3-17

Detail

BEGINNING WITH

\ '-...

· ·.......
'· ' 18262939
..

Detail J
REQUIRED ON
AUSTRALIAN AIRCRAFT

Figure 3-8.
3-18

Change 3

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

Detail I
REAR SEAT
BEGINNING WITH
AIRCRAFT SERIAL
18262940

Detail I
FRONT SEAT
BEGINNING WITH
AIRCRAFT SERIAL
18262940

'"'

Figure 3-8. Seat Belt and Shoulder Harness Installation(Sheet 3 of 3)

Detail K

INERTIA REEL
BEGINNING WITH
SERIAL
AIRCRAFT SERIAL
18262940
18262940

Figure 3-8. Seat Belt and Shoulder Harness Installation(Sheet 3 of 3)
Change 3

3-19

Detail A

1.
2.
3.
4.
5.

Headliner
Wire Bow
Zipper
Front Spar Shield
Stud

6.
7.
8.
9.

Overhead Skylight Moulding
Tiara
Coat Hook
Aft Upper Window Moulding

Figure 3-9.
3-39.

AUXILIARY.
a. FOLD-UP.

3-41. REMOVAL AND INSTALLATION.
a. Remove bolts securing seat structure to hinge
brackets.
b. Unsnap seat back from aft cabin wall. (1968 and
1969 Models).
c. Lift seat out.
d. Reverse preceding steps for reinstallation.
3-42. REPAIR. Replacement of defective parts is
recommended in repair of seats. However, a cracked framework may be welded, provided the crack is
not in an area of stress concentration (close to a
hinge or bearing point). The square-tube framework
is 6061 aluminum, heat-treated to a T-6 condition.
Use a heliarc weld on these seats, as torch welds
will destroy heat-treatment of frame structure. Figure 3-7 outlines instructions for replacing defective
cams on reclining seat backs.

3-20

CABIN UPHOLSTERY.
Change 2

Cabin Headliner Installation

3-40. DESCRIPTION. These seats are permanently
bolted to the cabin structure and have no adjustment
provisions. The seat structure is mounted on hinge
brackets with pivot bolts, thus allowing seat to be
pivoted upward to acquire more baggage area.

3-43.

Detail

Due to the wide selec-

tion of fabrics, styles and colors, it is impossible to
depict each particular type of upholstery. The following paragraphs describe general procedures which
will serve as a guide in removal and replacement of
upholstery. Major work, if possible, should be done
by an experienced mechanic. If the work must be
done by a mechanic unfamiliar with upholstery practices, the mechanic should make careful notes during
removal of each item to facilitate replacement later.
3-44. MATERIALS AND TOOLS. Materials and
tools will vary with the job. Scissors for trimming
upholstery to size and a dull-bladed putty knife for
wedging material beneath retainer strips are the
only tools required for most trim work. Use industrial rubber cement to hold soundproofing mats
and fabric edges in place. Refer to Section 18 for
thermo-plastic repairs.
3-45. SOUNDPROOFING. The aircraft is insulated
with spun glass mat-type insulation and a sound deadener compound applied to inner surfaces of skin in
most areas of cabin and baggage compartment. All
soundproofing material should be replaced in its
original position any time it is removed. A soundproofing panel is placed in gap getween wing and
fuselage and held in place by wing root fairings.
3-46.

CABIN HEADLINER.

(Refer to figure 3-9.)

3-47. REMOVAL AND INSTALLATION.
a. Detail A.
1. Remove sun visors, all inside finish strips
and plates, doorpost upper shields, front spar trim
shield, dome lights and any other visible retainers
securing
securing headliner.
headliner.
2. Work edges of headliner free from metal
teeth which hold fabric.
3. Starting at front of headliner, work headliner
down, removing screws through metal tabs which
hold wire bows to cabin top. Pry loose outer ends of
bows from retainers above doors. Detach each wire
bow in succession.
NOTE
Always work from front to rear when removing headliner.
4. Remove headliner assembly and bows from
aircraft,
NOTE
Due to difference in length and contour of
wire bows, each bow should be tagged to
assure proper location in headliner.
5.

Remove spun glass soundproofing panels.
NOTE

The lightweight soundproofing panels are
held in place with industrial rubber cement.
6. Reverse preceding steps for reinstallation.
Before installation, check all items concealed by
headliner for security. Use wide cloth tape to
secure loose wires to fuselage and to seal openings
in wing roots. Straighten tabs bent during removal
of headliner.
7. Apply cement to inside of skin in areas where
soundproofing panels are not supported by wire bows
and press soundproofing in place.
8. Insert wire bows into headliner seams and secure two bows at rear of headliner. Stretch material
along edges to properly center, but do not stretch it
tight enough to destroy ceiling contours or distort
wire bows. Secure edges of headliner with metal
teeth.
9. Work headliner forward, installing each
wire bow in place with metal tabs. Wedge ends of
wire bows into the retainer strips. Stretch headliner
just taut enough to avoid wrinkles and maintain a
smooth contour.
10. When all bows are in place and fabric edges
are secured, trim off excess fabric and reinstall
all items removed.
b. Detail B.
1. Remove sun visors, all inside finish strips
and plates, overhead console, upper doorpost shields
and any other visible retainers securing headliner.
2. Remove molding from fixed windows,
3. Remove screws securing headliner and carefully take down headliner.
4. Remove spun glass soundproofing panels

above headliner.
NOTE
The lightweight soundproofing panels are held
in place with industrial rubber cement.
5. Reverse preceding steps for reinstallation.
Before installation, check all items concealed by
headliner for security. Use wide cloth tape to secure
loose wires to fuselage and to seal openings in wing
roots.
3-48. UPHOLSTERY SIDE PANELS. Removal of
upholstery side panels is accomplished by removing
seats for access, then removing parts attaching
panels. Remove screws, retaining strips, arm
rests and ash trays as required to free panels. Automotive type spring clips attach most door panels. A
dull putty knife makes an excellent tool for prying
clips loose. When installing side panels, do not
over-tighten screws. Larger screws may be used
in enlarged holes as long as area behind hole is
checked for electrical wiring, fuel lines and other
components which might be damaged by using a longer screw.
3-49. CARPETING. Cabin area and baggage compartment carpeting is held in place by rubber cement,
small sheet metal screws and retaining strips through
the 1970 model aircraft. Beginning with 1971 model
aircraft the carpeting is secured by Velcro fasteners
for quick-removal and inspection. When fitting a
new carpet, use the old one as a pattern for trimming
and marking screw holes.
3-50.

SAFETY PROVISIONS.

3-51. CARGO TIE-DOWNS. Cargo tie-downs are
used to ensure baggage cannot enter seating area
during flight. Methods of attaching tie-downs are illustrated in figure 3-10. The eyebolt and nutplate can
be located at various points. The sliding tie-down lug
also utilizes the eyebolt and attaches to a seat rail.
A baggage net can be secured to the aft cabin wall and
floor for baggage security.
3-52. SAFETY BELTS. Safety belts should be replaced if frayed or cut, latches are defective or
stitching is broken. Attaching parts should be replaced if excessively worn or defective. The front
seat safety belts are attached to brackets bolted to
the cabin floor and the center seat safety belts are
attached to the seats themselves. The auxiliary seat
is provided with only one safety belt and is snapped
into clips bolted to the aircraft structure. Refer to
figure 3-8.
NOTE
Through 1970 model aircraft, when installing
front and center seat safety belts be sure the
belt half with the buckle is installed on the inboard side of the seat. Beginning with 1971
models the belt half with the buckle should be
installed on the outboard side of the seat to ensure proper operation of the shoulder harness.

Change 2

3-21

CARGO TIE-DOWN
DOWN RING

SEAT RAIL

THRU 18260825
Figure 3-10.

BEGINNING WITH 18260826

Cargo Tie-Down Rings

2. Mirror
3. Grommet
4. Nut
5. Washer

Figure 3-11.

Rear View Mirror Installation

3-53. SHOULDER HARNESS. Individual shoulder
harnesses may be installed for each seat except
auxiliary. Through 1970 model aircraft each harness
is attached to a clip bolted to the upper fuselage
structure. Beginning with 1971 model aircraft the
pilot and copilot harnesses are bolted to the upper
rear doorposts and the center seat harnesses are
bolted to the aft cabin structure. Component parts
should be replaced as outlined in the preceding
paragraph. Refer to figure 3-8. Beginning with aircraft 18262940, an inertia reel installation is offered.

SHOP NOTES:

3-22

Change 2

Refer to figure 3-8 for installation.
3-54. GLIDER TOW-HOOK. A glider tow-hook,
which is mounted in place of the tail tie-down ring,
is available for all models.
3-55. REAR VIEW MIRROR. A rear view mirror
may be installed on the cowl deck above instrument
panel. Figure 3-10 shows details for rear view
mirror installation.

SECTION 4
WINGS AND EMPENNAGE

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

4-1.

WINGS AND EMPENNAGE.

4-2.

WINGS.

Page
4-1
4-1
4-1
. 4-1
4-3
.
4-3
.4-3
4-3
4-3
4-3
4-3

Fin ..................
Description .............
Removal.
...............
Repair
...............
.
.
Installation
..........
Horizontal Stabilizer
Description .............
. . ..........
Removal
Repair . ............
installation
.............

NOTE

(See figure 4-1.)

4-3. DESCRIPTION. Each all-metal wing panel is
a semicantilever, semimonocoque type, with two
main spars and suitable ribs for the attachment of
the skin. Skin panels are riveted to ribs, spars and
stringers to complete the structure. An all-metal,
piano-hinged aileron, flap, and a detachable wing tip
are mounted on each wing assembly. A single,
rubberized, bladder-type fuel cell is mounted in the
inboard end of each wing. The leading edge of the
left wing may be equipped with landing and taxi lights,
Navigation/strobe lights are mounted at 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 maintenance stand when the
fastenings are loosened.
a. Remove wing root fairings and fairing plates.
b. Remove all wing inspection plates.
c. Drain fuel from cell of wing being removed.
d. Disconnect:
1. Electrical wires at wing root disconnects.
2. Fuel lines at wing root. (Observe precautions outlined in paragraph 12-3.)
3. Pitot line (left wing only) at wing root.
4. Wing leveler vacuum tube, if installed, at
wing root.
e. Slack off tension on aileron cables by loosening
turnbuckles, then disconnect cables at aileron bellcranks. Disconnect flap cables at turnbuckles above
headliner, and pull cables into wing root area.

........

4-3
4-3
4-3
4-4
4-4
4-4
4-4
. 4-4
4-4
4-4

To ease rerouting the cables, a guide wire
may be attached to each cable before it is
pulled free of the wing. Cable then may be
disconnected from wire. Leave guide wire
routed through the wing; it may be attached
again to the cable during reinstallation and
used to pull the cable into place.
f. Support wing at outboard end and disconnect
strut at wing fitting. (Refer to paragraph 4-10.) Tie
the strut up with wire to prevent it from swinging
down and straining strut-to-fuselage fitting. Loosen
lower strut fairing and slide fairing up the strut; the
strut may then be lowered without damage.
NOTE
It is recommended that flap be secured in
streamlined position with tape during wing
removal to prevent damage, since flap will
swing freely.
g. Mark position of wing attachment eccentric
bushings (Refer to figure 4-1); these bushings are
used to rig out "wing heaviness. "
h. Remove nuts, washers, bushings and bolts
attaching wing spars to fuselage.
NOTE
It may be necessary to rock the wings slightly
while pulling attaching bolts, or to use a long
drift punch to drive out attaching bolts.
i.

Remove wing and lay on padded stand.
Change 1

4-1

Detail

A
9

a'g

,^*^

\^^,

the fuel bay cover panels are of

NOTE
The forward bushing is approximately
half the length of the aft bushing.
*THRU AIRCRAFT SERIAL 18260825
AND A182-0136

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

Nut
Washer
Bolt
Bolt
Bushing
Washer

7.
8.
9.
10.
11.
12.
13.

Nut
Rub Strip
Moulding
Fairing
Screw
Inspection Plate
Flap

Figure 4-1.
4-2

Change 3

Wing Installation

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

Aileron
Wing Tip
Navigation/Strobe Light
Landing and Taxi Lights
Stall Warning Unit
Fuel Cell

4-5. REPAIR. A damaged wing panel may be repaired in accordance with instructions outlined in
Section 18. Extensive repairs of wing skin or structure are best accomplished using the wing repair jig,
which may be obtained from Cessna. The wing jig
serves not only as a holding fixture, making work on
the wing easier, but also assures the absolute alignment of the repaired wing.
4-6. INSTALLATION.
a. Hold wing in position and install bolts, bushings,
washers and nuts attaching wing spars to fuselage
fittings. Be sure eccentric bushings are positioned
as marked.
b. Install bolts, spacers and nuts to secure upper
and lower ends of wing strut to wing and fuselage fittings.
c. Route flap and aileron cables, using guide wires.
(Refer to note in paragraph 4-4.)
d. Connect:
1. Electric wires at wing root disconnects.
2. Fuel lines at wing root. (Observe precautions outlined in Section 12).
3. Pitot line (if left wing is being installed.)
4. Cabin ventilator hose at wing root.
5. Wing leveler vacuum tube, if installed, at
wing root.
e. Rig aileron system (Section 6).
f. Rig flap system (Section 7).
g. Refill wing fuel cell and check for leaks.
(Observe precautions outlined in Section 12).
h. Check operation of wing tip lights and landing
and taxi lights.
i. Check operation of fuel quantity indicator.
j. Install wing root fairings.
NOTE
Be sure to insert soundproofing panel in wing
gap, if such a panel was installed originally,
before replacing wing root fairings.
k. Install all wing inspection plates, interior panels
and upholstery.
4-7. ADJUSTMENT (CORRECTING "WING-HEAVY"
CONDITION).
(Refer to figure 4-1.) If considerable control wheel
pressure is required to keep the wings level in normal flight, a "wing-heavy" condition exists.
a. Remove wing fairing strip on the "wing-heavy"
side of the aircraft.
b. Loosen nut (7) and rotate bushings (5) simultaneously until the bushings are positioned with the
thick sides of the eccentrics up. This will lower the
trailing edge of the wing, and decrease "wing-heaviness" by increasing the angle-of-incidence of the
wing.
CAUTION
Be sure to rotate the eccentric bushings
simultaneously. Rotating them separately
will destroy the alignment between the
off-center bolt holes in the bushings, thus
exerting a shearing force on the bolt, with

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

WING STRUTS.

(See figure 4-2.)

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

FIN.

(See figure 4-3.)

4-13. DESCRIPTION. The vertical fin is primarily
of metal construction, consisting of ribs and spars
covered with skin. Fin tips are of ABS construction.
Hinge brackets at the fin rear spar attach the rudder.
4-14. REMOVAL. The vertical fin may be removed
without first removing the rudder. However, for
access and ease of handling, the rudder may be removed by following procedures outlined in Section 10.
a. Remove fairings on either side of fin.
b. Disconnect flashing beacon lead, tail navigation
light lead, antennas and antenna leads, and rudder
cables, if rudder has not been removed.
NOTE
The flashing beacon electrical lead that routes.
into the fuselage may be cut, then spliced (or
quick-disconnects used) at installation.
c. Remove screws attaching dorsal to fuselage.
d. Remove bolts attaching fin rear spar to fuselage
fitting.
e. Remove bolts attaching fin front spar to fuselage,
and remove fin.
Change 1

4-3

On AIRCRAFT SERIALS 18261960

NOTE
Beginning with serial 18263256, wrap strut with
Y-8562 polyurethane tape (3-M Co.) or equivalent
in the areas where strut fairings contact strut.
Locate tape splice at trailing edge of strut.

7.
Figure 4-2.
4-15. REPAIR. Fin repair should be accomplished
in accordance with applicable instructions outlined in
Section 18.
4-16. INSTALLATION. Reverse the procedures
outlined in paragraph 4-14 to install the vertical fin.
Be sure to check and reset rudder and elevator travel.
If any stop bolts were removed or settings disturbed,
the systems will have to be rigged. Refer to applicable sections in this manual for rigging procedures.
4-17.

HORIZONTAL STABILIZER.

(See figure 4-4.)

4-18. DESCRIPTION. The horizontal stabilizer is
primarily of all-metal construction, consisting of
ribs and spars covered with skin. Stabilizer tips
are of ABS construction. A formed metal leading edge is riveted to the assembly to complete the
structure. The elevator trim tab actuator is contained within the horizontal stabilizer. The underside of the stabilizer contains a covered opening which
provides access to the actuator. Hinges are located
on the rear spar assembly to support the elevators.

4-4

Change 3

13.

Tape

Wing Strut
4-19. 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 4-14.
c. Disconnect elevator trim control cables at cable
ends and turnbuckle inside tailcone. Remove stop
blocks, then remove pulleys which route the aft
cables into horizontal stabilizer. Pull cables out of
tailcone.
4-20. REPAIR. Horizontal stabilizer repair should
be accomplished in accordance with applicable instructions outlined in Section 18.
4-21. INSTALLATION. Reverse procedures outlined in paragraph 4-19 to install the horizontal
stabilizer. Rig elevator, elevator trim and rudder
systems as outlined in Sections 8, 9 and 10 consecutively. Check operation of tail navigation light and
flashing beacon.

THRU SERIALS 18261528
AND A182-0146

BEGINNING WITH SERIALS
18261529 AND A182-0147

1. Fin Assembly
2. Upper Rudder Hinge
4.
5.
6.
7.
8.
9.

Lower Rudder Hinge
Bolt
Washer
Nut
Bolt
Fairing

BEGINNING WITH SERIALS
18261529 AND A182-0147

Detail

Detail C

Detail B

NOTE
ttach Bolt Torques:
* 70-100 lb inches
140-225 lb inches

7
6 6
THRU SERIALS 18261528
AND A182-0146

Refer to Cessna Single Engine Service
Letters, SE72-3, February 11, 1972
and SE72-29, September 29, 1972 for
vertical fin inspection information.

Detail C
Figure 4-3.

Vertical Fin
Change 3

4-5

NOTE
Detail D

A kit is available from the
Cessna Service Parts Center for replacement of the
abrasion boots.

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

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

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

Figure 4-4.
4-6

Change 3

Horizontal Stabilizer

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

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

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
SECTION 5
LANDING GEAR AND BRAKES
TABLE OF CONTENTS

Page

LANDING GEAR ........................................ 5-1
D escription.............................................. 5-1
Main Landing Gear .................................... 5-2
Troubleshooting...................................... 5-2
Removal (Thru 18260825) ..................... 5-2
Removal (Beginning with 18260826)..... 5-2
Installation (Thru 18260826) .................. 5-3
Installation (Beginning with 18260826).. 5-3
Step Bracket Installation......................... 5-3
Brake Line Fairing Replacement ............ 5-3
Main Wheel Speed Fairing Removal
and Installation .................................. 5-7A
Removal of Tubular Strut Fairing ........... 5-7A
Main Wheel Removal ............................. 5-7
Disassembly (Cleveland) ............... 5-8
Inspection and Repair (Cleveland). 5-8
Assembly (Cleveland) .................... 5-8
Disassembly (McCauley) ............... 5-8
Inspection and Repair (McCauley). 5-8
Assembly (McCauley) .................... 5-8
Main Nose Wheel Thru-Bolt
Nut Torque Values .................... 5-10A
Main Wheel Installation ............................ 5-10A
Main Wheel Axle Removal ....................... 5-11
Main Wheel Axle Installation .................... 5-11
Main Wheel Alignment ............................. 5-12
Wheel Balancing ...................................... 5-12
5-12
N ose G ear ...................................................
Troubleshooting ....................................... 5-12
Removal and Installation .......................... 5-14
Nose Wheel Speed Fairing
Removal and Installation ..................... 5-14
Nose Wheel Removal and Installation..... 5-14A

Inspection and Repair
(Cleveland Wheel) ................... 5-16A
Assembly (Cleveland Wheel)........ 5-16A
Disassembly (McCauley Wheel)... 5-16A
Inspection and Repair (McCauley
Wheel) .................................... 5-17
Assembly (McCauley Wheel)........ 5-17
Wheel Balancing ................................. 5-17
Nose Gear Shock Strut ...................... 5-17
Disassembly.................................. 5-17
Assembly ...................................... 5-19
Torque Links........................................ 5-19
Shimmy Dampener.............................. 5-19
Nose Wheel Steering System ............. 5-20
Steering Bungee Assembly .......... 5-20
Nose Wheel Steering Adjustment. 5-20
BRA K ES ................................................... 5-22
Description........................................... 5-22
Troubleshooting................................... 5-22
Brake Master Cylinders....................... 5-23
Removal and Installation .............. 5-23
Repair ............................... .......... 5-23
Hydraulic Brake Lines ......................... 5-23
Wheel, Brake Assembly ...................... 5-23
Removal ........................................ 5-23
Inspection and Repair................... 5-23
Assembly ...................................... 5-23
Installation ..................................... 5-23
Check Brake Lining Thickness............ 5-23
Brake Lining Installation ..................... 5-23
Brake Bleeding .................................... 5-26
Parking Brake System........................ 5-26

Disassembly (Cleveland Wheel) .... 5-14A

5-1.

LANDING GEAR.

DESCRIPTION. Aircraft through Serial 18260825 are equipped with non-retractable, tricycle landing
5-2.
gear, utilizing flat spring-steel main gear struts. A bracket to attach a step to each strut is bonded to the main
landing gear spring-strut with a thermo-setting, high-strength cement. Beginning with aircraft Serial
18260826, these aircraft are equipped with tubular spring-steel main gear struts, also equipped with step
brackets. The main gear struts are enclosed by streamlined fairings. Wheel brake lines are routed through
the fairings to each main wheel. Disc-type brakes and tube-type tires are installed on the axle at the lower
end of the strut. Speed fairings or heavy-duty wheels may be installed on some aircraft. The nose gear is a
combination of a conventional air/oil (oleo) strut and fork, incorporating a shimmy dampener. The nose wheel
is steerable with the rudder pedals up to a maximum pedal deflection, after which it becomes free-swiveling
up to a maximum travel of 30 degrees right or left of center. Through the use of the brakes, the aircraft can
be pivoted about the outer wing strut fitting. A speed fairing or a heavy-duty shock strut and wheel may be
installed on some aircraft.
Revision 4
Mar 1/2004

5-1
© Cessna Aircraft Company

5-3.

MAIN LANDING GEAR.

5-4.

TROUBLE SHOOTING
TROUBLE

AIRCRAFT LEANS TO
ONE SIDE.

TIRES WEAR EXCESSIVELY.

WHEEL BOUNCE EVIDENT
EVEN ON SMOOTH SURFACE.

Incorrect tire inflation.

Inflate to correct pressure.

Loose or defective landing
gear attaching parts.

Tighten or install new parts.

Landing gear spring excessively
sprung.

Install new landing gear
spring-strut.

Incorrect shimming at inboard
end of spring-strut. (flat gear)

Install shims as required.

Bent axle(s).

Install new axle(s).

Incorrect tire inflation.

Inflate to correct pressure.

Wheels out of alignment.

Align wheels in accordance
with paragraph 5-19 and
figure 5-5.

Landing gear spring excessively
sprung.

Install new landing gear
spring-strut.

Incorrect shimming at
inboard end of spring-strut.

Install shims as required.
Refer to figure 5-1.

Bent axle(s).

Install new axle(s).

Dragging brakes.

Refer to paragraph 5-43.

Loose or defective wheel
bearings.

Adjust.
5-25.

Wheels out of balance.

Correct in accordance with
paragraph 5-20.

Out of balance condition.

Correct in accordance with
paragraph 5-20.

5-5. REMOVAL. (Thru 18260825, refer to figure
5-1, sheet 1.) This procedure removes the landing
gear as a complete assembly. Refer to applicable
paragraphs for removal of individual components.
a. Remove floorboard access covers over springstrut being removed.
b. Hoist or jack aircraft in accordance with paragraph 2-4 or 2-5.
c. Remove screws and slide external fairing plate
and seal down around spring-strut.
d. Drain hydraulic brake fluid from brake line on
spring-strut being removed.
e. Disconnect hydraulic brake line at bulkhead
fitting near inboard end of spring-strut so that brake
line is removed with the spring-strut. Cap or plug
disconnect fittings to prevent entry of foreign material into the fittings or line.
f. Remove channel at outboard forging by removing
5-2

REMEDY

PROBABLE CAUSE

See paragraphs 5-16 and

nuts, washers, and bolts.
g. Remove bolt attaching inboard end of springstrut to inboard forging and work entire gear out of
fuselage. Note shims placed under inboard end of
spring-strut and mark or tape shims together to be
sure they are installed correctly at installation of
the spring-strut.
5-6. REMOVAL. (Beginning with 18260826, refer
to figure 5-1, sheet 2.) This procedure removed the
landing gear as a complete assembly. Refer to applicable paragraphs for removal of individual components.
a. Jack or hoist aircraft as outlined in Section 2.
b. Remove brake bleeder screw and drain hydraulic
fluid from brake on gear being removed.
c. Remove screws from fuselage fairing and slide
down strut fairing for access to brake line.

d. Disconnect and cap or plug brake line at upper
end of strut.
e. Remove seats as necessary, peel back carpet
and remove access plates as necessary for access to
strut.
f. Remove snap ring (1) for strut-attaching pin (2).
g. Remove plug button (25) from belly of aircraft
below gear forging.
h. Using a punch, drive attaching pin upward out of
inboard fitting (26).
i. Pull strut outboard out of fittings (24) and (26).

j. Lower aircraft to ground.
k. Reinstall carpet and seats removed.
1. Check wheel alignment in accordance with
figure 5-4.
5-9.

STEP BRACKET INSTALLATION
NOTE
The step bracket is secured to the landing
gear spring strut with EA9309, or a similar
epoxy base adhesive. (Refer to figure 5-3.)

NOTE
To replace bushing from outboard fitting (24),
remove retaining ring at inboard end and
slide bushing outboard from forging. (Refer
to Section A-A.)
5-7. INSTALLATION. (Thru 18260825, refer to
figure 5-1, sheet 1. )
a. Slide landing gear fairing plate and seal over
upper end of landing gear spring-strut.
b. Slide spring-strut into place and work shims in
position under inboard end of spring-strut. Install
bolt, washer, and nut to secure inboard end of springstrut and shims to inboard forging,
NOTE
Shims (P/N 0541105) are installed under the
inboard end of the spring-strut to level the
wings within a tolerance of three inches.
Maximum number of shims permissible is
two.
c. Install channel at outboard forging with bolts,
washers, and nuts. Make sure arrow on channel
points outboard; it is possible to install channel incorrectly. Tighten channel attaching bolts evenly to
660-750 pound-inches with at least 80 per cent contact between channel and spring-strut. Also, tighten
inboard attach bolt to the correct torque for the size
bolt and nut. Torque chart for bolt and nuts sizes
are shown in figure 1-3.
d. Attach seal and external fairing with screws.
e. Lower aircraft and remove jack or hoist.
f. Connect hydraulic brake line; fill and bleed brake
system.
g. Install floorboard access covers and other components removed for access,
5-8. INSTALLATION. (Beginning with 18260826,
refer to figure 5-1, sheet 2.)
a. Reinstall all parts removed from strut.
b. Clean and polish machined surface on upper end
of strut. Prime fitting (10) per note, if required.
c. Apply Dow Corning Compound DC7 to unpainted
area on upper end of strut,
d. Slide strut through bushing into inboard forging
and align attaching pin holes.
e. Install attaching pin and snap ring.
f. Install access plates and plug button.
g. Remove caps or plugs and connect brake line.
h. Fill and bleed brake system in accordance with
paragraph 5-55.
i. Install fuselage fairing.

a. Mark position of the bracket so that the new step
bracket will be installed in approximately the same
position on the strut.
b. Remove all traces of the original bracket and
adhesive as well as any rust, paint or scale with
a wire brush and coarse sandpaper.
c. Leave surfaces slightly roughened or abraded,
but deep scratches or nicks should be avoided.
d. Clean surfaces to be bonded together thoroughly.
If a solvent is used, remove all traces of the solvent
with a clean, dry cloth. It is important that the bonding surfaces be clean and dry.
e. Check fit of the step bracket on the strut. A
small gap is permissible between bracket and strut.
f. Mix adhesive (EA9309) in accordance with manufacturer's directions.
g. Spread a coat of adhesive on bonding surfaces,
and place step bracket in position on strut. On the
flat spring-strut, tap the bracket upward on the strut
to ensure a good tight fit of the bracket on the strut.
On the tubular strut, clamp bracket to strut to ensure
a good tight fit.
h. Form a small fillet of the adhesive at all edges
of the bonded surfaces. Remove excess adhesive with
lacquer thinner.
i. Allow the adhesive to cure thoroughly according
to the manufacturer's recommendations before flexing
the strut or applying loads to the step.
j.
Paint the strut and step bracket after curing is
completed.
5-10. BRAKE LINE FAIRING REPLACEMENT.
(Refer to figure 5-1, sheet 1.)
a. Disconnect brake line at wheel and drain fluid,
or plug line to avoid draining. Flex brake line away.
b. Remove all traces of the original adhesive as
well as any rust, paint or scale with a wire brush and
coarse sandpaper. Sand inner surface of fairing strip,
running sanding marks lengthwise.
c. Leave surfaces slightly roughened or abraded.
Deep scratches or nicks should be avoided.
d. Clean surfaces to be bonded thoroughly. If a
solvent is used, remove all traces of the solvent with
a clean, dry cloth. It is important for the surfaces
to be clean and dry. Solvent should not be used on
the vinyl fairing strip.
e. Mix the adhesive (Saco 326 or Hysol EA-9311)
according to manufacturer's directions.
f. Apply a thin, uniform coat of adhesive to each
bonding surface. Pot life of Saco 326 is approximately 20 minutes at 77F. Pot life of Hysol EA9311 is approximately 5 minutes at 77°F. The
material will cure to 90% of its ultimate strength in
one hour, with complete cure in 24 hours.
5-3

NOTE

strut.

2.
3.

Inboard Forging
Outboard Forging

Clip

9.
10.
11.

Union
Hose Plate
BrakeCover
Disc

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

Shims
Axle
Brake
Bolt Assembly
Nut-4
Cotter Pin
Hub Cap

21.
22.

Step
Channel

Figure 5-1.
5-4

Main Landing Gear

(Sheet 1 of 2)

\ 'J ,

MAY BE USED AS AN ALTERNATE
THREAD LUBE ON THE PIPE
THREADED BRAKE FITTING ONLY.

COVER PLATE

13
OUTBOARD

FITTING
SPACER
RING

NOTE

surfaces only.
SECTION A-A

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

Apply Y8560 Polyurethane tape (3M Co., St. Paul, Minnesota)
to upper and lower surfaces of tubular strut (23) in area where
fairing (6) will cause chafing.

Ring
Pin
Upper Fairing
BrakeLine
Step Tread
Strut Fairing
Step Assembly
Hose
Bracket

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

Fitting
Shim
Axle
Torque Plate
Lower Fairing
Wheel Assembly
Axle Nut
Cotter Pin

18. Hub Cap
19. Back Plate
20. Brake Cylinder
21. Bracket
22. Sta-Strap
23. Strut Assembly
24. Outboard Fitting
25. Plug Button
26. Inboard Fitting

BEGINNING WITH 18260826

Figure 5-1.

Main Landing Gear

(Sheet 2 of 2)
Change 2

5-5

1971 THRU 1974 MODELS

PRIOR TO 1971 MODEL

1.
2.
3.

4.
5.
6.
7.

Mounting Plate
Speed Fairing
Scraper

Figure 5-2.
5-6

Change 2

Bolt
Axle Nut
Hub Cap
Stiffener

Main Wheel Fairings (Sheet 1 of 2)

8.
9.
10.

Doubler
Axle
Torque Plate

Detail A
BEGINNING WITH 1975 MODELS

Figure 5-2.

Main Wheel Fairings (Sheet 2 of 2)
Change 2

5-6A/(5-6B blank)

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
B3703

NOTE: After installation of screw (6),
cement entire forward half of
step tread (5) to step (7) with
EC880, EC1300 or equivalent.
(Refer to shaded area.)

1. Step Support Assembly
2. Sta-Strap
3. Spring Strut

4. Brake Line
5. Step Tread

Figure 5-3. Step Bracket Installation
G.

Position fairing strip between brake line and strut, and press firmly against strut. Press brake line
into groove of fairing strip and wrap immediately with masking tape in five equally-spaced places.
Excess adhesive may be removed with solvents.

Allow adhesive to cure thoroughly according to manufacturer's directions before flexing the gear.

After recommended curing time, remove tape and connect brake line.

Paint area as required.

Fill and bleed brake system.

5-11.
MAIN WHEEL SPEED FAIRING REMOVAL AND INSTALLATION. Main wheel speed fairings are
removed by removing the screws attaching the inboard side of the wheel speed fairing to the attach plate,
which is bolted to the axle, and removing the bolt securing the outboard side of the wheel speed fairing to the
axle nut. Loosen the scraper when necessary and work speed fairing from the wheel. Installation is the
reversal of the removal. After installation, check scraper-to-tire clearance for a minimum clearance of 0.56inch (9/16 inch) to a maximum of 0.69 inch (11/16 inch). Elongated holes in the scraper are provided so the
scraper can be adjusted. Refer to Service Kit SK182-12 for repair of the wheel speed fairings used on 19691970 model aircraft.
CAUTION:

Revision 4
Mar 1/2004

ALWAYS CHECK SCRAPER-TO-TIRE CLEARANCE AFTER INSTALLING SPEED
FAIRINGS, WHENEVER A TIRE HAS BEEN CHANGED AND WHENEVER SCRAPER
ADJUSTMENT HAS BEEN DISTURBED. IF THE AIRCRAFT IS FLOWN FROM
SURFACES WITH MUD, SNOW OR ICE, SPEED FAIRINGS SHOULD BE CHECKED TO
MAKE SURE THERE IS NO ACCUMULATION WHICH COULD PREVENT NORMAL
WHEEL ROTATION. WIPE FUEL AND OIL FROM SPEED FAIRINGS TO PREVENT
STAINS AND DETERIORATION.
5© Cessna Aircraft Company

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
5-11A.
A.

REMOVAL OF TUBULAR STRUT FAIRING. (Refer to Figure 5-1, Sheet 2.)
Remove 6 screws from perimeter and 3 screws from lower side of fuselage fairing (3).

B. Twist fuselage fairing and remove from strut fairing (6).
C. Repeat steps "A" and "B" for lower fairing (14) if speed fairings are installed.
D. Remove screws attaching step assembly (7); remove step assembly.
E.

Remove 9 screws from strut fairing (6).

Spread fairing (6) far enough to remove from strut (23).

G. If speed fairings are not installed, remove screws attaching cover plate.
H. Reverse preceding steps to install fairings.
5-12.

MAIN WHEEL REMOVAL. (See Figure 5-4.)
NOTE:

It is not necessary to remove the main wheel to reline brakes or remove brake parts, other
than the brake disc or torque plate.

5-7A/(5-7B Blank)
© Cessna Aircraft Company

Revision 4
Mar 1/2004

a. Hoist or jack aircraft as outlined in Section 2.
b. Remove speed fairing, if installed, in accordance with paragraph 5-11.
c. Remove hub cap, cotter pin, and axle nut.
d. Remove bolts and washers attaching back plate
to brake cylinder and remove back plate.
e. Pull wheel from axle.
5-13. MAIN WHEEL DISASSEMBLY (Cleveland).
a. Remove valve core and deflate tire. Break tire
beads loose from wheel rims.

WARNIN
Injury can result from attempting to separate
wheel halves with the tire inflated. Avoid
damaging wheel flanged when breaking tire
beads loose.
b. Remove thru-bolts and separate wheel halves,
removing tire and tube and brake disc.
c. Remove the grease seal rings, felts, and bearing cones from the wheel halves.
NOTE
The bearing cups are a press fit in the wheel
halves and should not be removed unless a
new part is to be installed. To remove the
bearing cups, heat wheel half in boiling water
for 30 minutes, or in an oven not to exceed
149°C (300°F). Using an arbor press, if
available, press out the bearing cup and press
in the new cup while the wheel is still hot.
5-14. INSPECTION AND REPAIR (Cleveland).
a. Clean all metal parts and grease seal felts in
cleaning solvent and dry thoroughly.
b. Inspect wheel halves for cracks. Cracked
wheel halves shall be discarded and new parts used.
Sand out small 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.
c. Inspect brake disc. If excessively warped,
scored, or worn to a thickness of 0. 190-inch, the
brake disc should be replaced with a new part. Sand
smooth small nicks and scratches.
d. Carefully inspect bearing cones and cups for
damage and discoloration. After cleaning, pack bearing cones with clean aircraft wheel bearing grease
(figure 2-5) before installing in wheel half.
5-15. MAIN WHEEL ASSEMBLY (Cleveland).
a. Insert thru-bolts through brake disc and position in the inner wheel half, using the bolts to guide
the disc. Ascertain that the disc is bottomed in the
wheel half.
b. Position tire and tube with the tube inflation
valve through hole in outboard wheel half.
c. Place the inner wheel half in position on outboard wheel half. Apply a light force to bring wheel
halves together. While maintaining the light force,
assemble a washer and nut on one thru-bolt and
tighten snugly. Assemble the remaining washers
and nuts on the thru-bolts and torque to the value
stipulated in figure 5-4A.
5 -8

Change 3

CAUTION
Uneven or improper torque of thru-bolt nuts
can cause failure of bolts, with resultant
wheel failure.
d. Clean and pack bearing cones with clean aircraft
bearing grease (figure 2-5).
e. Assemble bearing cones, grease seal felts, and
rings into wheel halves.
f. Inflate tire to seat tire beads, then adjust tire to
correct pressure.
5-15A. MAIN WHEEL DISASSEMBLY. (McCauley
Wheel.) (Refer to figure 5-4.)
a. Remove valve core and deflate tire and tube.
Break tire beads loose from wheel flanges.

WARNING
Injury can result from attempting to remove
wheel flanges with the tire and tube inflated.
Avoid damaging wheel flanges when breaking
tire beads loose.
b. Remove thru bolts, nuts and washers or capscrews and washers (whichever are installed. )
c. Separate wheel flanges from wheel hub. Retain
spacers between wheel flanges and wheel hub.
d. Remove wheel hub from tire and tube.
e. Remove retainer rings, grease seal retainers,
grease seal felts and bearing cones from wheel hub.
NOTE
The bearing cups 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 hub is still
hot.
5-15B. MAIN WHEEL INSPECTION AND REPAIR.
(McCauley Wheel )
a. Clean all metal parts, grease seal felts and mylar spacers in cleaning solvent and dry thoroughly.
b. Inspect wheel flanges and wheel hub for cracks.
Discard cracked wheel flanges or hub and install new
parts. Sand out nicks, gouges and corroded areas.
When protective coating has been removed, clean the
area thoroughly, prime with zinc chromate and paint
with aluminum lacquer.
c. If excessively warped or scored, or worn to a
thickness of 0. 190-inch, brake disc should be replaced with a new part. Sand smooth small nicks and
scratches.
d. Carefully inspect bearing cones and cups for
damage and discoloration. After cleaning, pack bearing cones with clean aircraft wheel bearing grease
(refer to Section 2) before installing in the wheel hub.
5-15C. MAIN WHEEL REASSEMBLY. (McCauley
Wheel )
a. Place wheel hub in tire and tube with tube inflation

17
1.
2.
3.
4.
5.
6.
7.

Snap Ring
Grease Seal Ring
Grease Seal Felt
Grease Seal Ring
Bearing Cone
Outer Wheel Half
Tire

8.
9.
10.
11.
12.
13.
14.
15.
Figure 5-4.

Tube
Inner Wheel Half
Bearing Cup
Brake Disc
Bushing
Torque Plate
Pressure Plate
Anchor Bolt

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

Brake Cylinder
Brake Bleeder
O-Ring
Piston
Thru-Bolt
Brake Lining
Back Plate

Wheel and Brake Assembly (Sheet 1 of 2)
Change 1

5-9

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

Snap Ring/
Grease Seal Retainer (Outboard)
Grease Seal Felt (Outboard)
Grease Seal Retainer (Outboard)
Bearing Cone
Wheel Flange (Aluminum)
Spacer
Tire
Tube
Wheel Hub
Bearing Cup (Race)
Grease Seal Retainer (Inboard)
Grease Seal Felt (Inboard)
Brake Disc
Torque Plate
Pressure Plate

Figure 5-4.
5-10

Change 3

22
/
20
1
17. Anchor Bolt
18. Brake Cylinder
19. Bleeder Valve
20. O-Ring (Piston)
21. Brake Piston
22. Brake Lining
23. Thru-Bolt
24. Back Plate
25. Capscrew
26. Wheel Flange (Steel)

Wheel and Brake Assembly

(Sheet 2 of 2)

MAIN
GEAR

NOSE
GEAR

WHEEL NUMBER

SIZE

1241156-11

6.00 X 6

1241156-12

NUT/ CAP
SCREW TORQUE

FLANGE

CLEVELAND

150 lb-in

MAGNESIUM

5.00 X 5

CLEVELAND

90 lb-in

MAGNESIUM

C163002-0201

5.00 X 5

MC CAULEY

90-100 Ib-in

C163003-0201

5.00 X 5

MC CAULEY

90-100 Ib-in

STEEL

C163003-0401

5.00 X 5

MC CAULEY

*190-200 lb-in

STEEL

C163001-0103

6.00 X 6

CLEVELAND

150 lb-in

MAGNESIUM

C163001-0104

6.00 X 6

CLEVELAND

90 lb-in

ALUMINUM

C163002-0101

6.00 X 6

MC CAULEY

90-100 lb-in

ALUMINUM

C163003-0102

6.00 X 6

MC CAULEY

*190-200 lb-in

Figure 5-4A.

MANUFACTURER

ALUMINUM

STEEL

Main and Nose Wheel Thru-Bolt Nut or Capscrew Torque Values

stem in cutout of wheel hub.
b. Place spacer and wheel flange on inboard side of
wheel hub (opposite of tube inflation stem).
c. Place washer under head of each thru-bolt and
insert bolt through wheel flange and wheel hub, or
place washer under head of each capscrew and start
capscrews into wheel hub threads.
d. Place spacer and wheel flange on other side and
align valve stem in cutout in wheel flange.
e. Install washers and nuts on thru-bolts or place
washer under head of each capscrew and start capscrews into wheel hub threads.

WARNING
Be sure that spacers and wheel flanges are
seated on flange of wheel hub. Uneven or
improper torque of thru-bolt nuts or capscrews can cause failure of the bolts or capscrews, with resultant wheel failure.
f. Tighten thru-bolt nuts or capscrews evenly and
torque to values specified in figure 5-2A.
g. Clean and pack bearing cones with clean aircraft
wheel bearing grease. (Refer to Section 2 for grease
type.)
h. Assemble bearing cones, grease seal felts and
retainer into wheel hub.
i. Inflate tire to seat tire beads, then adjust to correct tire pressure. Refer to figure 1-1 for correct
tire pressure.
5-15D. MAIN AND NOSE WHEEL THRU-BOLT NUT
OR CAPSCREW TORQUE VALUES. (Refer to figure
5-4A. ) During assembly of the main or nose wheel,
thru-bolt nuts or capscrews should be tightened evenly and torqued to the values specified in figure 5-4A.

To facilitate identification of wheel manufacturers,
solid wheels are manufactured by Cleveland Products
Co. , and webbed wheels are manufactured by
McCauley Industrial Corporation. Cleveland wheels
are also identified by having two wheel halves as
shown in figure 5-4 (sheet 1 of 2) and figure 5-7.
McCauley wheels are identified by having two wheel
flanges and a hub as illustrated in figure 5-4 (sheet
2 of 2) and figure 5-7. The differences between
McCauley steel-flange and aluminum-flange wheels
are illustrated in figures 5-4 (sheet 2 of 2) and figure
5-7.
5-16.

MAIN WHEEL INSTALLATION.

a. Place wheel assembly on axle.
b. Install axle nut and tighten nut 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.
d. Install hub cap. Install speed fairing (if used)
as outlined in paragraph 5-11.
CAUTION
Always check scraper-to-tire clearance after
installing speed fairings, whenever a tire has
been changed, and whenever scraper adjustment has been disturbed. If the aircraft is
flown from surfaces with mud, snow, or ice,
the fairing should be checked to make sure
there is no accumulation which could prevent
normal wheel rotation. Refer to paragraph
5-9 for correct scraper-to-tire clearance.

Change 3

5-10A

ALUMINUM PLATES, APPROXIMATELY
18" SQUARE, PLACED UNDER WHEELS
GREASE BETWEEN PLATES
NOTE
checking wheel alignment.

TOP VIEW

FRONT VIEW

OF TOE-IN CHECK

OF CAMBER CHECK

Measure camber by reading protractor level
held vertically against outboard flanges of
wheel.

Measure toe-in at edges of wheel flange. Difference in measurements is toe-in for one wheel.
(half of total toe-in. )

POSITIVE CAMBER 7

CARPENTER'S SQUARE

FORWARD

-NEGATIVE CAMBER

INBOARD

STRAIGHTEDGE
NOTE
Setting toe-in and camber within these tolerances while the cabin and fuel tanks are empty will give
approximately zero toe-in and zero camber at gross weight. Therefore, if normal operation is at
less than gross weight and abnormal tire wear occurs, realign the wheels to attain the ideal setting
for the load conditions. Refer to sheet 2 of this figure for shims availability and their usage. Always use the least number of shims possible to obtain the desired result.
Figure 5-5.
5-10B

Change 2

Main Wheel Alignment (Sheet 1 of 2)

CORRECTION IMPOSED ON WHEEL

SHIM
PART
NO.

POSITION OF
THICKEST CORNER
OR EDGE OF SHIM

TOE-IN

TOE-OUT

POS. CAMBER

NEG. CAMBER

0541157-1

AFT
FWD

.06"
----

---.06"

---0°3 '

003'

0541157-2

UP
DOWN

.006"
----

---.006"

0030
----

1241061-1

UP & FWD
UP & AFT
DOWN & FWD
DOWN & AFT

.03"
.06"
-------

------.06"
.03"

2050
2049
-------

'
'

--2049 '
2050'

0441139-5

UP & FWD
UP & AFT
DOWN & FWD
DOWN & AFT

---.12"
---.11"

.11"
---.12"
----

0025
0°11
-------

----0°11 '
0025'

0441139-6

UP & FWD
UP & AFT
DOWN& FWD
DOWN & AFT

---.24"
---.22"

.22"
---.24"
----

0050
0022
-------

0541157-3

AFT
FWD

.12"
----

---.12"

---007 '

--0030

-----0022 '
0050'

007
---

1241061-1
0541157-3
0541157-2
0541157-1
0441139-6
0441139-5
1241061-1
0541157-3
0541157-2
0541157-1
0441139-6
0441139-5
SHIM NO.

COLUMN 1

Figure 5-5.

0
0
0
0
0
0

0
2
2
2
1
1

0
1
2
2
1
2

0
0
1
1
0
0

0
1
1
1
0
1

Max. number of
shims to be used
with shims in
column 1.
COLUMN 2

Main Wheel Alignment (Sheet 2 of 2)

5-17. MAIN WHEEL AXLE REMOVAL.
a. Remove speed fairing in accordance with paragraph 5-11.
b. Remove wheel in accordance with paragraph 5-12.
c. Disconnect, drain, and plug or cap the hydraulic
brake line at the wheel brake cylinder.
d. Remove nuts, washers and bolts securing axle,
brake components and speed fairing mounting plate,
if used, to strut (flat gear) or strut-attach fitting
(tubular gear).

NOTE
When removing axle from strut or strut-attach
fitting, note number and position of wheel
alignment shims between axle and strut or
attach fitting. Mark shims or tape together
carefully so they can be installed in exactly
the same position to ensure wheel alignment
is not disturbed.
5-18. MAIN WHEEL AXLE INSTALLATION.
a. Secure axle and brake components to strut or
Change 3

5-11

strut-attach fitting, assuring that wheel alignment
shims and speed fairing mounting plate, if used, are
installed in their original positions.
b. Install wheel assembly on axle in accordance
with paragraph 5-16.
c. Connect hydraulic brake line to wheel brake
cylinder.
d. Fill and bleed affected brake system in accordance with paragraph 5-55.
e. Install speed fairing, if used, in accordance
with paragraph 5-11.
5-19. MAIN WHEEL ALIGNMENT. Correct main
wheel alignment is obtained through the use of tapered shims between the flange of the axle and spring
strut. See figure 5-5 for procedure to use in wheel
alignment. Wheel shims and the correction imposed
on the wheel by the various shims are listed in the
illustration.

5-21.

NOSE GEAR.

5-22.

TROUBLE SHOOTING.
TROUBLE

NOTE
Failure to obtain acceptable wheel alignment
through the use of the shims indicate a deformed main gear spring-strut or strut attaching bulkhead out of alignment.
5-20. WHEEL BALANCING. Since uneven tire wear
is usually the cause of wheel unbalance, replacing
the tire probably will correct this condition. Tire
and tube manufacturing tolerances permit a specified
amount of static unbalance. The light-weight point
of the tire is marked with a red dot on the tire sidewall and the heavy-weight point of the tube is marked
with a contrasting color line (usually near the inflation 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 the Cessna Service Parts Center.

PROBABLE CAUSE

REMEDY

TIRES WEAR EXCESSIVELY.

Loose nose gear torque links.

Check looseness and add shims
as required or install new parts.
See figure 5-10.

NOSE WHEEL SHIMMY.

Nose gear strut attaching clamps
loose.

Tighten nose gear strut attaching
clamp bolts.

Shimmy dampener needs fluid.

Service in accordance with
Section 2.

Defective shimmy dampener.

Repair or install new shimmy
dampener.

Loose or worn nose wheel
steering linkage.

Tighten loose linkage or replace
defective parts.

HYDRAULIC FLUID LEAKAGE FROM NOSE GEAR STRUT.

Defective nose gear strut seals
or defective parts.

Strut overhaul in accordance with
paragraphs 5-34 and 5-35.

NOSE GEAR STRUT WILL
NOT HOLD AIR PRESSURE.

Defective air filler valve or
valve not tight.

Check gasket and tighten loose
valve. Install new valve if
defective.

Defective nose gear strut
seals.

Install new seals.
5-34 and 5-35.

5-12

Change 3

See paragraphs

BEGINNING
WITH 18261426

NOTE
.shaded parts of the nose gear turn
the nose gear steering system is
erated on the ground, but do not turn
ile airborne. As the lower strut extends, a centering block on the upper
torque link contacts a flat spot on the
ttom end of the upper strut, thus

keeping
the lower strut

MAXIMUM EXTENSION
(Thru 18260825)
(Beginning with 18260826)

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

Bolt
Nut
Upper Forging
Bolt
Upper Strut
Steering Bungee
Lower Forging

and wheel

from

turning.
5.00" ±. 15"
4. 85" ±. 15"

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

Upper Torque Link
Bolt
Lower Torque Link
Torque Link Fitting
Nose Gear Fork
Wheel and Tire
Bolt

15.
16.
17.
18.
19.
20.
21.
22.

Figure 5-6.

Bolt
Steering Collar
Screw
Bolt
Steering Torque Arm
Shimmy Dampener
Bolt
Closure Assembly

Nose Gear Installation
Change '

5-13

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
5-23.

REMOVAL AND INSTALLATION.

Remove engine cowling for access.

B. Weight or tie down the tail to raise nose wheel off ground.
C.

Disconnect bungee and shimmy dampener from nose gear.

D. Remove air filler valve core and deflate strut completely and telescope strut to its shortest length.
WARNING:

BE SURE THE STRUT IS DEFLATED COMPLETELY BEFORE REMOVING BOLT
AT TOP OF STRUT.

Remove bolt through upper forging and strut.

Either of two methods may be used to remove the strut from the aircraft. The following procedure
outlines removing the strut along with the lower forging at the fuselage. An alternate method is to
remove and disconnect parts as required to slide the strut down through the lower forging, leaving
the forging attached to the fuselage.

Remove four bolts attaching lower forging to fuselage. Remove rudder bar shields from inside
the cabin for access to the nuts.

Pull strut assembly down, out of upper forging to remove.

Installation of the nose gear strut is the reversal of the preceding steps. Always install bolt at top
forging before clamping strut in lower forging to prevent misalignment.
NOSE WHEEL SPEED FAIRING REMOVAL AND INSTALLATION.

5-24.
A.

Weight or tie down tail of aircraft to raise nose wheel off the floor.

Remove nose wheel axle stud.
WARNING:

DEFLATE STRUT BEFORE REMOVING BOLT ATTACHING COVER PLATE,
FAIRING, AND TOW BAR SPACERS.

C. Deflate strut and remove bolts securing cover plate, fairing, and tow bar spacers to strut. Remove
cover plate.
D. Slide speed fairing up and remove nose wheel. Loosen scraper as necessary.
G. Install speed fairing by reversing the preceding steps. Tighten axle stud until a slight bearing drag is
obvious when the wheel is rotated. Back off nut to the nearest castellation and install cotter pins.
H. Service shock strut after installation has been completed.
CAUTION:

ALWAYS CHECK SCRAPER CLEARANCE AFTER INSTALLING SPEED FAIRING,
WHENEVER A TIRE HAS BEEN CHANGED AND WHENEVER SCRAPER
ADJUSTMENT HAS BEEN DISTURBED. SET CLEARANCE BETWEEN TIRE AND
SCRAPER TO A MINIMUM OF 0.56 INCH (9/16 INCH) TO A MAXIMUM OF 0.69
INCH (11/16 INCH). ELONGATED HOLES IN THE SCRAPER ARE PROVIDED
FOR ADJUSTMENT. IF THE AIRCRAFT IS FLOWN FROM SURFACES WITH MUD,
SNOW OR ICE, SPEED FAIRINGS SHOULD BE CHECKED TO MAKE SURE
THERE IS NO ACCUMULATION WHICH COULD PREVENT NORMAL WHEEL
ROTATION. WIPE FUEL AND OIL FROM SPEED FAIRINGS TO PREVENT STAINS
AND DETERIORATION.

5-14
© Cessna Aircraft Company

Revision 4
Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 SKYLANE SERIES
SERVICE MANUAL
5-25.

NOSE WHEEL REMOVAL AND INSTALLATION.

Weight or tie down tail of aircraft to raise the nose wheel off the ground.

Remove nose wheel axle bolt.

Pull nose wheel assembly from fork and remove spacers and axle tube from nose wheel. Loosen
scraper if necessary.

Reverse the preceding steps to install nose wheel. Tighten axle bolt until a slight bearing drag is
obvious when wheel is rotated. Back the nut off to the nearest castellation and install cotter pin.
CAUTION:

5-26.

ON AIRCRAFT EQUIPPED WITH SPEED FAIRINGS, ALWAYS CHECK SCRAPERTO-TIRE CLEARANCE AFTER INSTALLING SPEED FAIRING, WHENEVER A TIRE
HAS BEEN CHANGED, OR WHENEVER SCRAPER ADJUSTMENT HAS BEEN
DISTURBED. SET SCRAPER CLEARANCE INACCORDANCE WITH PARAGRAPH
5-24.

NOSE WHEEL DISASSEMBLY (Cleveland Wheel).
A.

Remove hub cap, completely deflate tire and break tire beads loose.
WARNING:

INJURY CAN RESULT FROM ATTEMPTING TO SEPARATE WHEEL HALVES
WITH THE TIRE INFLATED. AVOID DAMAGING WHEEL FLANGES WHEN
BREAKING TIRE BEADS LOOSE.

B. Remove thru-bolts and separate wheel halves.
C.

Remove tire and tube from wheel halves.

D. Remove bearing retaining rings, grease felt seals and bearing cones.
NOTE:

Revision 4
Mar 1/2004

The bearing cups are a press-fit in the wheel halves and should not be removed unless a
new part is to be installed. To remove the bearing cups, heat wheel half in boiling water for
30 minutes, or in an oven not to exceed 149°C (300°F). Using an arbor press, if available,
press out the bearing cup and press in the new cup while the wheel is still hot.

5-14A/(5-14B Blank)
© Cessna Aircraft Company

CLEVELAND NOSE WHEEL

Figure 5-7.

1. Snap Ring
2. Grease Seal Ring
3. Bearing
4. Tire
5. Tube
6. Grease Seal Felt
7. Thru-Bolt

Nose Wheels

8. Bearing Cup
9. Male Wheel Half
10. Female Wheel Half
11. Washer
12. Nut
13. Retaining Ring
14. Grease Seal Ring
15. Bearing

Figure 5-7.

McCAULEY NOSE WHEEL

16. Wheel Flange (Aluminum)
17. Spacer
18. Tire
19. Tube
20. Hub Assembly
21. Thru-Bolt
22. Grease Seal Felt
23. Wheel Flange (Steel)
24. Capscrew

Nose Wheels
Change 3

5-15

1971 THRU 1974 MODELS

1. Speed Fairing
2. Tow-Bar Spacer
3. Cover Plate

Figure 5-8.
5-16

Change 2

4.
5.
6.

Fork Bolt
Scraper
Axle Stud

Nose Wheel Speed Fairings (Sheet 1 of 2)

7.
8.
9.

Ferrule
Hub Cap
Access Door

BEGINNING WITH 1975 MODELS

Figure 5-8.

Nose Wheel Speed Fairings (Sheet 2 of 2)

5-27. NOSE WHEEL INSPECTION AND REPAIR
(Cleveland Wheel).
Instructions outlined in paragraph 5-14 for the main
wheel may be used as a guide for inspection and repair of the nose wheel.
5-28. NOSE WHEEL ASSEMBLY (Cleveland Wheel).
a. Insert tire and tube on wheel half and position
valve stem through hole in wheel half.
b. Insert thru-bolts, position other wheel half, and
secure with nuts and washers. Take care to avoid
pinching tube between wheel halves. Tighten bolts
evenly to torque value stipulated in figure 5-4A.

d. Assemble bearing cones, seals, and retainers
into the wheel halves.
e. Inflate tire to seat tire beads, then adjust to
correct pressure.
f. Install spacers, axle tube and hub cups, and
install wheel assembly in accordance with paragraph
5-25.
5-29. NOSE WHEEL DISASSEMBLY. (McCauley
Wheel.)
a. Remove hub caps, completely deflate tire and
break tire beads loose at wheel flanges.

WARNING
Uneven or improper torque on the thru-bolt
nuts may cause bolt failure with resultant
wheel failure.
c. Clean and pack bearing cones with clean aircraft wheel bearing grease (figure 2-5).

wheel flanges with tire and tube inflated.
Avoid damaging wheel flanges when breaking
tire beads loose.
b. Remove thru-bolt nut, washers and thru-bolts
or capscrews and washers.
Change 3

5-16A/(5-16B blank)

c. Separate wheel flanges from wheel hub. Retain
spacers between wheel flanges and wheel hub.
d. Remove wheel hub from tire and tube.
e. Remove retainer rings and remove grease seal
retainers, grease seal felts and bearing cones from
wheel hub.
NOTE
The bearing cups 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 hub is still
hot.
5-30. NOSE WHEEL INSPECTION AND REPAIR
(McCauley Wheel.)
a. Clean all metal parts, grease seal felts and mylar spacers in cleaning solvent and dry thoroughly.
b. Inspect wheel flanges and wheel hub for cracks.
Cracked wheel flanges or hubs shall be discarded and
new parts installed. Sand out smooth all 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.
c. Carefully inspect bearing cones and cups for
damage and discoloration. After cleaning, pack bearing cones with clean aircraft wheel bearing grease before installing in the wheel hub. (Refer to Section 2
for grease type. )
5-31. NOSE WHEEL ASSEMBLY. (McCauley
Wheel.)
a. Insert tube in tire, aligning index marks on tire
and tube.
b. Place wheel hub in tire with valve stem in cutout
of wheel hub.
c. Place spacer and wheel flange on one side of hub.
d. Place washer under head of each thru-bolt and
insert bolt through wheel flange and wheel hub, or
place washer under head of each capscrew and start
capscrews into wheel hub threads.

e. Place spacer and wheel flange on other side and
align valve stem in cutout in wheel flange.
f. Install washers and nuts on thru-bolts, or place
washer under head of each capscrew and start capscrews into wheel hub threads.
CAUTION
Be sure that spacers and wheel flanges are
seated on flange of wheel hub. Uneven or
improper torque of thru-bolt nuts or capscrews can cause failure of the bolts or capscrews with resultant wheel failure.
g. Tighten thru-bolts or capscrews evenly and torque to the values specified in figure 5-2A.
h. Clean and pack bearing cones with clean aircraft
grease. (Refer to Section 2 for grease type.)
i. Assemble bearing cones, grease seal felts and
retainer into wheel hub.
j. Inflate tire to seat tire beads, then adjust to correct tire pressure. (Refer to Section 1.)
5-32. WHEEL BALANCING. Refer to paragraph
5-20 for wheel balancing information.
5-33. NOSE GEAR SHOCK STRUT. (Refer to figure
5-8.) Removal and installation of the nose gear is
accomplished as outlined in paragraph 5-23. Speed
fairing and wheel removal and installation information is outlined in paragraph 5-24 and 5-25. The
heavy-duty nose gear is illustrated in figure 5-11
which may be used as a guide during maintenance.
Removal, installation, disassembly and assembly
procedures are the same as those outlined for the
standard nose gear strut except for the differences
illustrated in figure 5-11.
5-34. NOSE GEAR SHOCK STRUT DISASSEMBLY.
(Refer to figure 5-8.) This procedure applies to
disassembly of the nose gear shock strut after it has
been removed from the aircraft, and the speed fairing and nose wheel have been removed. In many
cases, separation of the upper and lower strut will
permit inspection and parts installation without removal or complete disassembly of the strut.

SHOP NOTES:

Change 3

5-17

NOTE
Shims are available to use
as required above washer (10).

2
13

13.

Lower Strut

14.

Packing Support Ring

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

Scraper Ring
Retaining Ring
Lock Ring
Nut
Nut
Metering Pin
0-RPacking
Base Plug
Nut
O-Ring

Valve
0-Ring
3. Orifice Piston Support
4. Upper Strut
5. Decal
6. Steering Torque Arm
7. Screw
8. Retaining Ring
9. Steering Collar
10. Washer
11. Lock Ring
12. Bearing

Figure 5-9.
5-18

Change 1

25.
26.
27.
28.
29.
30.
31.
32.

Nose Gear Shock Strut

Fork
Bolt
Bolt
Toxque Link Fitting
Back-Up Ring
0-Ring
0-Ring
Closure Assembly*

WARNING
~

c. Sharp metal edges should be smoothed with No.
400 emery paper, then thoroughly cleaned with sol-

Be sure strut is completely deflated before

vent.

removing lock ring in lower end of upper
strut, or disconnecting torque links.

d. Used sparingly, Dow Corning DC-4 compound is
recommended for O-ring lubrication. All other internal parts should be liberally coated with hydraulic
fluid during assembly.

a. Remove shimmy dampener.
b. Remove torque links. Note position of washers,
shims and spacers.
c. Remove steering torque arm and lower forging
if these items have not been removed previously.
d. Remove lock ring from groove inside lower end
of upper strut. A small hole is provided at the lock
ring groove to facilitate removal of the lock ring.
NOTE
Hydraulic fluid will drain from strut as lower
strut is pulled from upper strut.
e. Use a straight, sharp pull to separate upper and
lower struts. Invert lower strut and drain remaining
hydraulic fluid.
f. Remove lock ring and bearing at top of lower
strut.
g. Slide packing support ring, scraper ring, retaining ring, and lock ring from lower strut, noting
relative position and top side of each ring; wire together if desired.
h. Remove O-rings and back-up rings from packing support ring.
i. Remove attaching torque link fitting and remove
torque link fitting from lower strut.
NOTE
Bolt attaching torque link fitting also holds
metering pin base plug in place.
j. Push metering pin and base plug assembly from
lower strut. Remove O-rings and metering pin from
base plug.
NOTE
Lower strut and fork are a press fit, drilled
on assembly. Separation of these parts is
not recommended, except for installation of
new part.

NOTE
Cleanliness and proper lubrication, along with
careful workmanship are important during
assembly of the shock strut.
e. When installing steering torque arm, lubricate
needle bearing in torque arm with general purpose
grease (figure 2-5) before installing. If needle bearing is defective, install new steering torque arm
assembly. Use shims as required between steering
torque arm and washer to provide a snug fit with retainer ring installed. Shims are available from the
Cessna Service Parts Center as follows:
1243030-5 ..........
1243030-6 ..........
1243030-7 ..........

0.006 inch
0.012 inch
0.020 inch

f. When installing lock ring in lower end of upper
strut, position lock ring so that one of its ends covers
the small access hole in the lock ring groove at the
bottom of the upper strut.
g. Temporary bolts or pins of correct diameter and
length are useful tools for holding parts in correct
relation to each other during assembly and installation.
h. After assembly of strut, install in accordance
with paragraph 5-33.
i. After installation, service shock strut as outlined in Section 2.
5-36. TORQUE LINKS. (Refer to figure 5-10.) The
illustration may be used as a guide during disassembly and assembly. The torque links keep the lower
strut aligned with the nose gear steering system, but
permit shock strut action. Torque link bushings
should not be removed except for replacement with
new parts. Excessively worn parts should be replaced with new parts.

WARNING
k. Remove retaining ring securing steering arm
assembly on upper strut and remove steering arm,
shims, and washer.
1. Push orifice support from upper strut and remove O-ring.
5-35. NOSE GEAR SHOCK STRUT ASSEMBLY.
(Refer to figure 5-9.)
a. Thoroughly clean all parts in cleaning solvent
and inspect them carefully. Replace all worn or defective parts and all O-rings and back-up rings with
new parts.
b. Assemble the strut by reversing the order of the
procedure outlined in paragraph 5-35 with the exception that special attention must be paid to the following
procedures.

Always deflate nose gear strut before disconnecting torque links.
5-37. SHIMMY DAMPENER. The shimmy dampener
offers resistance to shimmy by forcing hydraulic
fluid through small orifices in a piston. The dampener piston shaft is secured to a stationary part, and
the housing is secured to the nose wheel steering
torque arm assembly, which moves as the nose wheel
is turned, causing relative motion between the dampener shaft and housing. The shimmy dampener is illustrated in figure 5-12, which may be used as a guide
during disassembly and assembly. When assembling
the dampener, use new O-rings. Lubricate parts
with clean hydraulic fluid during assembly. Refer to
Section 2 for servicing procedures.
5-19

5-38. NOSE WHEEL STEERING SYSTEM. Nose
wheel steering is accomplished through the use of
the rudder pedals. A steering bungee links the nose
gear to a whiffletree which is operated by push-pull
rods connected to the rudder bars. Steering is
afforded up to approximately 10 degrees each side of
center, after which brakes may be used to gain a
maximum deflection of 30 degrees right or left of
center. A flexible boot is used to seal the fuselage
entrance of the steering bungee. A sprocket-operated screw mechanism to provide rudder trim is incorporated at the aft end of the bungee. Refer to
Section 10 for the rudder trim system.

5-39. STEERING BUNGEE ASSEMBLY. The bungee
assembly is spring-loaded and should not be disassembled internally. The steering bungee is connected to
the steering torque arm on the strut by a bearing end
assembly and to the steering whiffletree by a rod end
assembly.
5-40. NOSE WHEEL STEERING ADJUSTMENT.
Since the nose wheel steering, rudder system, and
rudder trim systems are interconnected, adjustments to one system will affect the others. Section
10 contains rigging instructions for the nose wheel
steering system as well as the rudder and rudder
trim systems.

NOTE
Tighten bolts (8) to 20-25 poundinches, then safety the bolts by
bending tips of safety lug (10).

Refer to figure 5-9 for
remainder of strut.

Tighten nuts (7) snugly, then
tighten to align next castellation
with cotter pin hole.
Shims (3) are available to use as
required to remove any looseness.

1.
2.
3.
4.
5.

Spacer
Grease Fitting
Shim
Bushing
Stop Lug

Figure 5-10.
5-20

6.
7.
8.
9.
10.

Upper Torque Link
Nut
Bolt
Lower Torque Link
Safety Lug

Torque Links

1.
2.
3.

Hub
Lower Strut
Tow-Bar Spacer

Figure 5-11.

4.
5.
6.

Lugs
Fork
Bushing

Heavy-Duty Shock Strut

THREAD INSERT

NOTE
Orifice in piston (10) connects
to passage in rod (7).

RIFICE

13
1/16" HOLE

NOTE
When installing the shimmy dampener,
use washers as required between the
dampener and the steering torque arm
to cause a snug fit.

1. Retainer
2. O-Ring
3. Bearing Head
4. Barrel

5.
6.
7.
8.
9.

Figure 5-12.

Stat-O-Seal
Filler Plug
Rod
Back-Up Ring
Roll Pin

10.
11.
12.
13.

Piston
Floating Piston
Spring
Set Screw

Shimmy Dampener
5-21

5-41.

BRAKE SYSTEM.

5-42. DESCRIPTION. The hydraulic brake system
is comprised of two master brake cylinders, located
immediately forward of the rudder pedals, brake

5-43.

TROUBLE SHOOTING.

TROUBLE
DRAGGING BRAKES.

BRAKES FAIL TO
OPERATE.

5-22

lines connecting each master cylinder to its wheel
brake cylinder, and the single-disc, floating cylindertype brake assembly, located at each main landing
gear wheel

PROBABLE CAUSE

REMEDY

Brake pedal binding.

Lubricate pivot points; replace
or repair defective parts.

Weak or broken piston return
spring in master cylinder.

Repair or replace master cylinder.

Parking brake control improperly
adjusted.

Adjust properly.

Insufficient clearance between
lock-O-seal and piston in
master cylinder.

Adjust clearance per figure 5-13.

Restriction in hydraulic lines or in
passage in master cylinder
compensating sleeve.

Remove restrictions; flush brake
system with denatured alcohol.
Repair or replace master cylinder.

Warped or badly scored brake
disc.

Replace disc and linings.

Damage or accumulated dirt
restricting free movement of
wheel brakes.

Clean and repair or replace brake
parts.

Fluid low in master cylinder
or wheel cylinder.

Fill system and bleed brakes.

Faulty O-rings in master cylinder
or wheel cylinder.

Replace O-rings.

Faulty lock-O-seal in master
cylinder.

Replace lock-O-seal.

Excessive clearance between lockO-seal and piston.

Adjust clearance per figure 5-13.

Internal damage to hose and O-rings
due to use of wrong type of hydraulic fluid.

Replace damaged parts. Flush
system with denatured alcohol. Fill
and bleed brake system.

Pressure leak in system.

Tighten connection; repair or
replace faulty parts.

Brake linings worn out.

Replace linings.

Oil or grease on brake linings or
new linings just installed.

Clean linings with carbon tetrachloride.

5-44. BRAKE MASTER CYLINDERS. The brake
master cylinders, located just forward of the pilot
rudder pedals, are actuated by applying pressure at
the top of the rudder pedals. A small reservoir is
incorporated into each master cylinder to supply it
with fluid. When dual brakes are installed, mechanical linkage permits the copilot pedals to operate the
master cylinders.
5-45. REMOVAL AND INSTALLATION.
a. Remove bleeder screw at wheel brake assembly
and drain hydraulic fluid from brake cylinder.
b. Remove front seats and rudder bar shield for
access to the brake master cylinders,
c. Disconnect parking brake linkage and disconnect
brake master cylinders from rudder pedals.
d. Disconnect brake master cylinders at lower attach points.
e. Disconnect hydraulic hose from brake master
cylinders and remove cylinders.
f. Plug or cap hydraulic fittings, hose, and lines
to prevent entry of foreign matter.
g. Reverse the preceding steps to install brake master cylinders, then fill and bleed brake system in
accordance with paragraph 5-55.
5-46. REPAIR. (Refer to figure 5-13.) Cylinder
breakdown is shown in the figure which may be used
as a guide during disassembly, adjustment and assembly. Repair is limited to installation of new parts,
cleaning and adjustment. During assembly, use clean
hydraulic fluid as a lubricant.
5-47. HYDRAULIC BRAKE LINES. The lines are of
rigid tubing, except for flexible hose used at the
brake master cylinders and at the wheel cylinders on
the flat spring strut equipped aircraft. A separate
line is used to connect each brake master cylinder
to its corresponding wheel brake cylinder.
5-48. WHEEL BRAKE ASSEMBLY. (Refer to figure
5-3.) The wheel brake assemblies use a disc which
is attached to the main wheel with the wheel thrubolts. The brake assemblies are also equipped with
a floating brake assembly.
5-49. REMOVAL. (Refer to figure 5-1.) Wheel
brake assemblies are the floating type and can be
removed after disconnecting the brake hose and
removing the back plate.
NOTE
The brake disc is removed after wheel
removal and disassembly. To remove
the torque plate, remove wheel and
axle as outlined in paragraph 5-17.
5-50. INSPECTION AND REPAIR.
a. Clean all parts except brake linings and O-rings
in dry cleaning solvent and dry thoroughly.
b. New O-rings are usually installed at each overhaul. If O-ring re-use is necessary, they should be
wiped with a clean cloth soaked in hydraulic fluid and
inspected for damage.

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 maximum permissible wear. See paragraph 5-53.
d. Inspect brake cylinder wall for scoring. A
scored cylinder will leak or cause rapid O-ring wear.
Install new brake cylinder.
e. If the anchor bolts on 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 drive new bolts in with a soft
mallet.
f. Inspect wheel brake disc for a minimum thickness of 0. 190-inch. If brake disc is below minimum
thickness, warped or out of round, install a new part.
5-51. ASSEMBLY. (Refer to figure 5-4.) The
figure may be used as a guide during assembly.
Lubricate parts with clean hydraulic fluid and assemble parts with care to prevent damage to O-rings.
5-52. INSTALLATION. Place brake assembly in
position with pressure plate in place, then install
back plate. If torque plate was removed, install as
the axle is installed. If the brake disc was removed
from the wheel, install as wheel is assembled.
5-53. CHECKING BRAKE LINING THICKNESS.
Lining should be replaced if worn to a minimum
thickness of 3/32-inch. Visually compare a 3/32inch strip of material held adjacent to each lining to
measure thickness of the lining. The shank end of
the correct size drill bit makes an excellent tool for
checking minimum thickness of brake linings.
5-54. BRAKE LINING INSTALLATION. (Refer to
figure 5-4. )
a. Remove bolts securing back plate and remove
back plate.
b. Pull the 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 and pressure plate in the same manner.
NOTE
A rivet setting kit, Part No. R561, is available from the Cessna Service Parts Center.
This kit consists of an anvil and punch.
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
rivet head against the anvil.
f. Center the rivet setting punch on the lips of the
rivet. While holding the back plate down firmly

5-23

NOTE

pressure cannot build up in the
reservoir during brake operation.
Remove plug and drill 1/16" hole,
30 ° from vertical, if plug is not

0. 040 ±0. 005 INCH

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

Clevis
Jamb Nut
Piston Rod
Cover
Setscrew
Cover Boss

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

Body
Reservoir
O-Ring
Cylinder
Piston Return Spring
Nut

Figure 5-13.
5-24

Brake Master Cylinder

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

Piston Spring
Piston
Lock-O-Seal
Compensating Sleeve
Filler Plug
Screw

18260669 & A182-0136

Detail A

1. Attaching Angle
2. Stiffener Angle
3. Handle

5. Clamp
6. Cotter Pin
7. Positioning Pin
8. Cable Assembly
9. Brake Master Cylinder
10. Brake Line
11. Brake Hose
12. Bracket
13. Bellcrank
14. Cable
15. Pin
16. Spring
17. Pulley

17
Detail B

Figure 5-14.

Parking Brake System (Sheet 1 of 2)
5-25

(REFER TO SHEET 1)
ROUTING BEGINNING

WITH 18260670 & A182-0137
12

Detail

BEGINNING WITH
18260670 & A182-0137

(REFER TO SHEET 1)

Figure 5-14.

Parking Brake System (Sheet 2 of 2)

against the lining, hit the punch with a hammer to
set the rivet. Repeat blows on the punch until lining
is firmly against the back plate.
g. Realign the lining on the back plate and install
rivets in 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.
5-55. BRAKE BLEEDING. Standard bleeding, with
a clean hydraulic pressure source connected to the
wheel cylinder bleeder, is recommended.
a. Remove brake master cylinder filler plug and
screw flexible hose with appropriate fitting into the
filler hole at top of master cylinder. Immerse the
free end of the hose in a container with enough hydraulic fluid to cover the end of the hose.
b. Connect a clean hydraulic pressure source, such
as a hydraulic hand pump or Hydro Fill unit, to the
5-26

bleeder valve in the wheel cylinder.
c. As fluid is pumped into the system, observe the
immersed end of the hose at the brake 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.
NOTE
Ensure that the free end of the hose from the
master cylinder remains immersed during the
entire bleeding process.
5-56. PARKING BRAKE SYSTEM. (Refer to figure
5-14. )
The parking brake system uses a handle and ratchet
mechanism connected by a cable to linkage at the
master cylinders. Pulling out on the handle depresses
both cylinder piston rods and the ratchet locks the
handle in this position until handle is turned and released.

SECTION 6
AILERON CONTROL SYSTEM

TABLE OF CONTENTS

Page

AILERON CONTROL SYSTEM .........
Description ...............
Trouble Shooting ...........
..
Control Column .............
Description .............
Removal and Installation .......
Pilot's Control Column ......
Copilot's Control Column .....
Repair. ..
...
. ...
. ..
...
Aileron Bellcrank ............

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

DESCRIPTION.

6-3.

TROUBLE SHOOTING.

6-1
6-1
6-1
6-2
6-2
6-2
6-2
6-6
6- 6
6-6

(Refer to fig-

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

.......

. ...

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

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

The aileron control system is

NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 6-18.
TROUBLE
LOST MOTION IN CONTROL
WHEEL.

RESISTANCE TO CONTROL
WHEEL MOVEMENT.

PROBABLE CAUSE

REMEDY

Loose control cables.

Check cable tension. Adjust
cables to proper tension.

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

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

Cables too tight.

Check cable tension. Adjust
cables to proper tension.

Pulleys binding or cable off.

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

Bellcrank distorted or
damaged.

Check visually.
bellcrank.

Replace defective

Defective quadrant assembly.

Check visually.
quadrant.

Replace defective

Clevis bolts in system too
tight.

Check connections where used.
Loosen, then tighten properly
and safety.
Change 2

6-1

6-3.

TROUBLE SHOOTING (Cont).

TROUBLE
CONTROL WHEELS NOT
LEVEL WITH AILERONS
NEUTRAL.

PROBABLE CAUSE

REMEDY

Improper adjustment of
cables.

Refer to paragraph 6-18.

Improper adjustment of
aileron push-pull rods.

Adjust push-pull rods to obtain
proper alignment.

DUAL CONTROL WHEELS
NOT COORDINATED.

Cables improperly adjusted.

Refer to paragraph 6-18.

INCORRECT AILERON
TRAVEL.

Push-pull rods not adjusted
properly.

Refer to paragraph 6-18.

Incorrect adjustment of travel
stop bolts.

Refer to paragraph 6-18.

6-4.

CONTROL COLUMN.

(Refer to figure 6-2.)

6-5. DESCRIPTION. Rotation of the control wheel
rotates four bearing roller assemblies (3) on the end
of the control wheel tube (13), which in turn, rotates
a square control tube assembly (15) inside and extending from the control wheel tube (13). Attached
to this square tube (15) is a quadrant (24) 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 (29), turnbuckle (30) and adjustment terminals
(27). The forward end of the square control tube (15)
is mounted in a bearing block (21) on firewall (31) and
does not move fore-and-aft, but rotates with the control wheel. The four bearing roller assemblies (3)
on the end of the control wheel tube reduce friction
as the control wheel is moved fore-and-aft for elevator system operation. A sleeve weld assembly (5),
containing bearings which permit the control wheel
tube to rotate within it, is secured to the control
wheel tube by a sleeve and retaining ring in such a
manner it moves fore-and-aft with the control wheel
tube. This movement allows the push-pull tube (16)
attached to the sleeve weld assembly (5) to operate
an elevator arm assembly (18), to which one elevator
cable (20) is attached. A torque tube (19) connects
this arm assembly (18) to the one on the opposite end
of the torque tube (19), 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. PILOT'S CONTROL COLUMN.
1. (THRU AIRCRAFT SERIAL 18260825.) (Refer to figure 6-2, sheet 1.) Remove screws attaching
control wheel (2) to control wheel tube assembly (13)
and remove wheel. Disconnect electrical wiring to
map light and mike switch, if installed.
6-2

2. (BEGINNING WITH AIRCRAFT SERIAL 18260826. ) (Refer to figure 6-2, sheet 2.) Slide cover
(2) toward instrument panel to expose adapter (3).
Remove screws securing adapter (3) to control wheel
tube assembly (1) and remove control wheel assembly.
Disconnect electrical wiring to map light and mike
switch at connector (17), if installed. Slide cover (2)
off control wheel tube assembly (1).
3. (Refer to figure 6-2, sheet 1.) Remove decorative cover from instrument panel.
4. Remove screw securing adjustable glide plug
(14) to control tube assembly (15) and remove plug
and glide assembly.
5. Disconnect push-pull tube (16) at sleeve weld
assembly (5).
6. Remove screws securing support plate (10) at
instrument panel.
NOTE
To ease removal of control wheel tube assembly (13), snap ring (9) may be removed from
its locking groove to allow sleeve weld assembly (5) additional movement.
7. Using care, pull control wheel tube assembly
(13) aft and work assembly out through instrument
panel.
NOTE
If removal of control tube assembly (15) or
quadrant (24) is necessary, proceed to step
8.
8. Remove safety wire and relieve direct cable
tension at turnbuckles (index 8, figure 6-1).
9. Remove safety wire and relieve interconnect
cable tension at turnbuckle (30).
10. Remove safety wire and remove roll pin (28)
through quadrant (24) and control tube assembly (15).

Detail

....

B
Detail C

Detail

D
NOTE

1.
2.
3.
4.
5.

Cable Guard
Pulley
Spacer
Bushing
Rub Strip
NOTE (Car40
RubSri

Figure 6-1.

MAINTAIN PROPER CONTROL
CABLE TENSION.

ABL TENSION:
LBS ± 10 LBS ON AILERON CARRY-

Aileron Control System
6-3

Bearing

Thrust Bearing

10.
11.
12.
13.

Support Plate
Spacer
Collar
Control Wheel Tube

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

Control Tube Assembly
Push-Pull Tube
Support
Arm Assembly
Elevator Torque Tube
Elevator Control Cable

22.

Support

25.
26.

Nut
Idler Shaft

28.
29.
30.
31.

Adjustment Terminal
Roll Pin
Interconnect Cable
Interconnect Cable
Turnbuckle
Firewall

33.

Retainer

13
3 15 33
14

.PER SIDE

<,
NOTE
* Used only on aircraft
equipped with single
controls.

Washers (32) are of
various thicknesses

NOTE
31

Allow 0.030 " maximum
clearance between bearaft

tightening.
tter

Figure 6-2.
6-4

Change 1

ASERIALS 18260446 AND A1820137

18260825 AND A182-0136

Control Column Installation (Sheet 1 of 2)

AIRCRAFT SERIALS 18260826
THRU 18263475

* Plug (12) is used when mike

3.00
AIRCRAFT SERIALS
18263476 AND ON

NOTE
Torque bolt (19) to 30 lb-inches

Figure 6-2.

3. Adapter
4. Rubber
5. Plate
6. Map Light Rheostat
7. Terminal Block
8. Map Light Assembly
9. Control Wheel
10. Pad

13. Insulator
14. Plug
15. Bracket
16. Cable
17. Connector
18. Screw
19 Bolt

Control Column Installation (Sheet 2 of 2)
Change 3

6-5

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

Hinge
Balance Weight
Aileron
Pivot Bolt
Turnbuckle (Carry-Thru)
Bolt
Stop Bushing

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

Bellcrank
Turnbuckle (Direct)
Bushing
Brass Washer
Push-Pull Rod
Needle Bearing
Bushing

Figure 6-3.

Change 2

Brass washers (11) may be used as required
between lower end of bellcrank and wing channel to shim out excess clearance.

Aileron Installation

11. Remove pin, nut (25) and washer from control tube assembly (15) protruding through bearing
block (21) on forward side of firewall (31).
12. Using care, pull control tube assembly (15)
aft and remove quadrant (24).
13. Reverse the preceding steps for reinstallation. Safety wire all items previously safetied,
check rigging of aileron and elevator control systems
and rig, if necessary, in accordance with paragraph
6-18 and 8-14 respectively.
b. COPILOT'S CONTROL COLUMN.
1. Complete steps 1, 2, 3, 5, 6, 8, 9, 10 and
11 of subparagraph "a."
2. Using care, pull control tube assemblies
(13 and 15) aft and remove quadrant (24).
3. Remove radios, radio dust covers, cooling
pans and associated equipment as necessary to work
control wheel tube assembly (13) out from under instrument panel.
4. Complete step 13 of subparagraph "a. "

6-6

Carry-thru cable turnbuckle (5) may be located at either the right or left aileron bellcrank.

6-7. REPAIR. Worn, damaged or defective shafts,
bearings, drums, cables or other components should
be replaced. Refer to Section 2 for lubrication requirements.
6-8.

AILERON BELLCRANK.

(Refer to figure 6-3.)

6-9. REMOVAL.
a. Remove access plate inboard of each bellcrank
(8) on underside of wing.
b. Remove safety wire and relieve cable tension at
turnbuckle (5).
c. Disconnect control cables from bellcrank (8).
Retain all spacers and bushings.
d. Disconnect push-pull rod (12) at bellcrank.
e. Remove nuts, washers and bolts securing bellcrank stop bushing (7) and bellcrank (8) to wing structure.
f. Remove bellcrank through access opening, using
care that bushing (14) is not dropped from bellcrank.

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

Figure 6-4.

Inclinometer for Measuring Control Surface Travel
d. Remove cable guards and pulleys as necessary
to work cables free of aircraft.

NOTE
Brass washers (11) may be used as shims
between lower end of bellcrank and wing
structure. Retain these shims. Tape
open ends of bellcrank to prevent dust and
dirt from entering bellcrank needle bearings
(13).
6-10. REPAIR. Repair of bellcranks consists of
replacement of defective parts. If needle bearings
are dirty or in need of lubrication, clean thoroughly
and lubricate as outlined in Section 2.
6-11. INSTALLATION.
a. Place bushing (14) and stop bushing (7) in bellcrank (8) and position bellcrank in wing.
b. Install brass washers (11) between lower end of
bellcrank (8) and wing structure to shim out excess
clearance.
c. Install bellcrank pivot bolt (4).
d. Position bellcrank stop-bushing (7) and install
attaching bolt (6).
e. Connect control cables to bellcrank.
f. Connect push-pull rod (12) to bellcrank.
g. Re-rig aileron system in accordance with paragraph 6-18, safety turnbuckle (5) and reinstall all
items removed for access.
6-12.
6-1.)

CABLES AND PULLEYS.

(Refer to figure

6-13. REMOVAL AND INSTALLATION.
a. Remove access plates, wing root fairings and
upholstery as required.
b. Remove safety wire and relieve cable tension at
turnbuckles (8).
c. Disconnect cables from aileron bellcranks (7)
and quadrants (index 24, figure 6-2).

NOTE
To ease routing of cables, a length of wire
may be attached to end of cable before
being withdrawn from aircraft. Leave wire
in place, routed through structure; then
attach cable being installed and use to pull
cable into position.
e. After cable is routed in position, install pulleys
and cable guards. Ensure cable is positioned in pulley groove before installing guard.
f. Re-rig aileron system in accordance with paragraph 6-18, safety turnbuckles and install access
plates, fairings and upholstery removed in step "a."
6-14.

AILERONS.

(Refer to figure 6-3.)

6-15. REMOVAL.
a. Disconnect push-pull rod (12) at aileron.
b. Remove screws and nuts attaching aileron hinges
(1) to trailing edge of wing.
c. Using care, pull aileron out and down to slide
hinges from under wing skin and auxiliary spar reinforcements.
6-16. INSTALLATION.
a. Position aileron hinges between skin and auxiliary spar reinforcements and install screws and nuts
attaching hinges to trailing edge of wing.
b. Attach push-pull rod (12) to aileron.
NOTE
If rigging was correct and push-pull rod
adjustment was not disturbed, it should
not be necessary to re-rig system.
Change 2

6-7

c. Check aileron travel and alignment, re-rig if necessary, in accordance with paragraph 6-18.
6-17. REPAIR. Aileron repair may be accomplished
in accordance with instructions outlined in Section 18.
Before installation, ensure balance weights and hinges
are securely attached.
6-18. RIGGING. (Refer to figure 6-1.)
a. Remove safety wire and relieve cable tension at
turnbuckles (6 and 8).
b. Disconnect push-pull rods at bellcranks (7).
c. Adjust interconnect cable turnbuckle (index 30,
figure 6-2) and adjustment terminals (index 27, figure 6-2) to remove cable slack, acquire proper tension (40 ± 10 pounds) and position control wheels
level (synchronized).
d. Tape a bar across both control wheels to hold
them in neutral positon.
e. Adjust direct cable turnbuckles (8) and carrythru cable turnbuckle (6) so bellcrank stop-bushings
(index 7, figure 6-3) are centered in both bellcrank

6-8

Change 2

slots with 40±10 pounds tension on carry-thru cable.
Disregard tension on direct cables.
f. Adjust push-pull rods (index 12, figure 6-3) at
each aileron until ailerons are neutral with reference
to trailing edge of wing flaps. Be sure wing flaps
are full UP when making this adjustment.
g. With ailerons in neutral position (streamlined),
mount an inclinometer on trailing edge of one aileron
and set to 0° . (Refer to figure 6-4 for inclinometer.)
h. Remove bar from control wheels and check degree of travel as specified in figure 1-1. If travel is
not within specified limits, readjust push-pull rods
and cables as necessary.
i. Ensure all turnbuckles are safetied, all cables
and cable guards are properly installed, all jam nuts
are tight and replace all items removed for access.

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

SECTION 7
WING FLAP CONTROL SYSTEM

Page

TABLE OF CONTENTS
...
WING FLAP CONTROL SYSTEM . ...
...
..........
Description .
. ..
...
Operational Check . ....
.....
.... ...
Trouble Shooting .
Flap Motor and Transmission Assembly
Removal and Installation . .....
.............
Repair .
....
......
Flap Control Lever .
Removal and Installation ........
Drive Pulley ....
. . ........

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

7-1
7-1
7-1
7-2
.7-4

7-4
7-4
7-4
7-4
7-8

(Refer to

7-2. DESCRIPTION. The wing flap control system
consists of an electric motor and transmission assembly, drive pulleys, push-pull rods, cables, pulleys and follow-up control. Power from the motor
and transmission assembly is transmitted to the
flaps by a system of drive pulleys and cables. Electrical power to the motor is controlled by two microswitches mounted on a "floating" arm, a camming
lever and a follow-up control. As the camming lever
is moved to the desired flap setting, it trips a switch
actuating the flap motor. As the flaps move, the
floating arm is rotated by the follow-up control until
the active switch clears the camming lever, breaking
the circuit. To reverse direction of travel, the control lever is moved in the opposite direction. When
its cam contacts the second switch it reverses the
flap motor. Likewise the follow-up control moves
the floating arm until the second switch is clear of
the camming lever. Limit switches at the drive pulley are connected in series with the switches on the
floating arm to prevent over-travel of the flaps in the
full UP or DOWN position.
7-3. OPERATIONAL CHECK.
a. Operate flaps through their full range of travel,
observing for uneven or jumpy motion, binding and
lost motion in system. Ensure flaps are moving together through their full range of travel.
b. AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN NOT MODI-

7-8
Removal and Installation ........
. 7-8
Repair .............
. . . ..
7-8
. ....
....
Flaps
..
. 7-8
Removal and Installation .......
7-8
........
Repair ...
7-8
Cables and Pulleys ..............
7-8
. ...
Removal and Installation ..
. 7-8
Rigging - Flaps ...........
Rigging - Flap Control Lever and
7-13
Follow-Up ..........

FIED IN ACCORDANCE WITH FIGURE 7-2, SHEET
3. Attempt to overrun travel extremes and check for
transmission free-wheeling at full up and full down
positions.
c. BEGINNING WITH AIRCRAFT SERIALS 18260699
AND A182-0137 AND ALL AIRCRAFT MODIFIED IN
ACCORDANCE WITH FIGURE 7-2, SHEET 3. Operate flaps and check up-limit and down-limit switch
actuation in their respective positions.
d. Check that flaps are not sluggish in operation.
In flight at 110 mph, indicated airspeed, flaps should
fully extend in approximately 16.6 seconds and retract
in approximately 7. 0 seconds. On the ground, with
engine running, the flaps should extend in approximately 9.2 seconds and retract in approximately 8.2 seconds.
e. With flaps full UP, mount an inclinometer on one
flap and set to 0 ° . Lower flaps to full DOWN position
and check flap angle as specified in figure 1-1. Check
mid-range percentage setting, (approximate), against
degrees as indicated on inclinometer. Repeat the
same procedure for opposite flap.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
f. Remove access plates and attempt to rock drive
pulleys to check for bearing wear.
g. Inspect flap rollers and tracks for evidence of
binding and defective parts.
7-1

7-4.

TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraphs 7-18 and 7-19.
TROUBLE

BOTH FLAPS FAIL TO MOVE.

BINDING IN SYSTEM AS FLAPS
ARE RAISED AND LOWERED.

LEFT FLAP FAILS TO MOVE.

FLAPS FAIL TO RETRACT.

7-2

PROBABLE CAUSE

REMEDY

Popped circuit breaker.

Reset and check continuity.
Replace breaker if defective.

Defective switch.

Place jumper across switch.
Replace switch if defective.

Defective motor.

Remove and bench test.
Replace motor if defective.

Broken or disconnected wires.

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

Disconnected or defective
transmission.

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

Defective limit switch.

Check continuity of switches.
Replace switches found defective.

Follow-up control disconnected or slipping,

Secure control or replace
if defective.

Cables not riding on pulleys.

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

Bind in drive pulleys.

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

Broken or binding pulleys.

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

Frayed cable.

Check condition of cables.
defective cables.

Flaps binding on tracks.

Observe flap tracks and rollers.
Replace defective parts.

Disconnected or broken cable.

Check cable tension.
Connect or replace cable.

Disconnected push-pull rod.

Attach push-pull rod.

Disconnected or defective
UP limit switch.

Check continuity of switch.
Connect or replace switch.

Replace

Detail

FIGURE 7-4

_it

NOTE
Shaded pulleys are used
for this system only.
1.

Bushing

2.
3.

Pulley
Bracket

Spacer

9.
10.

Flap

MAINTAIN PROPER CONTROL

Rub Strip
Turnbuckle

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

Figure 7-1.

Wing Flap Control System
7-3

7-4.

TROUBLE SHOOTING (Cont).
PROBABLE CAUSE

TROUBLE

REMEDY

FLAPS FAIL TO EXTEND.

Disconnected or defective
DOWN limit switch.

Check continuity of switch.
Connect or replace switch.

INCORRECT FLAP TRAVEL.

Incorrect rigging.

Refer to paragraph 7-18.

Defective limit switch.

Check continuity of switches.
Replace switches found defective.

7-5. FLAP MOTOR AND TRANSMISSION ASSEMBLY.
7-6. REMOVAL AND INSTALLATION.
a. AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN NOT MODIFIED IN ACCORDANCE WITH SK150-37 AND WHEN
NOT MODIFIED IN ACCORDANCE WITH FIGURE
7-2, SHEET 3. (Refer to figure 7-2, sheet 1.)
1. Run flaps to full DOWN position.
2. Disconnect battery cables at the battery and
insulate cable terminals as a safety precaution.
3. Remove access plates adjacent to drive pulley and motor assembly on right wing.
NOTE
Remove motor (1), transmission (4), hinge
assembly (2) and actuating tube (8) from
aircraft as a unit on aircraft equipped with
standard fuel cells. On aircraft equipped
with long range cells, detach motor and
transmission assembly from hinge assembly (2) prior to removal.
4. Remove bolt (18) securing actuating tube (8)
to drive pulley (17).
5. Screw actuating tube (8) IN toward transmission (4) by hand to its shortest length.
6. Remove bolt (3) securing flap motor hinge
(2) to wing, or if long range fuel cells are installed,
remove bolt (5) securing transmission to hinge assembly. Retain brass washer between hinge and wing
structure for use on reinstallation.
7. Disconnect motor electrical wiring (21) at
quick-disconnects.
8. Using care, work assembly from wing through
access opening.
9. Reverse the preceding steps for reinstallation.
If the hinge assembly (2) was removed from the transmission (4) for any reason, ensure the short end of
hinge is reinstalled toward the top.
10. Complete an operational check as outlined in
paragraph 7-3 and re-rig system in accordance with
paragraph 7-18.

7-4

b. AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN MODIFIED
IN ACCORDANCE WITH SK150-37 AND WHEN NOT
MODIFIED IN ACCORDANCE WITH FIGURE 7-2,
SHEET 3. (Refer to figure 7-2, sheet 2.)
1. Complete steps 1, 3 and 4 of subparagraph
"a."
2. Run flap motor to place actuating tube (8) IN
to its shortest length.
3. Complete steps 2, 6, 7, 8, 9 and 10 of subparagraph "a."
c. BEGINNING WITH AIRCRAFT SERIALS 18260699 AND A182-0137 AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH FIGURE 7-2, SHEET
3. (Refer to figure 7-2, sheet 2.)
1. Complete steps 1 thru 7 of subparagraph "a."
2. Disconnect electrical wiring at limit switches
(29 and 32).
3. Complete steps 8, 9 and 10 of subparagraph
"a."
7-7. REPAIR. Repair consists of replacement of
motor, transmission, coupling, actuating tube and
associated hardware. Bearing in hinge assembly
may also be replaced. Lubricate as outlined in
Section 2.
7-8.

FLAP CONTROL LEVER.

7-9. REMOVAL AND INSTALLATION.
a. THRU AIRCRAFT SERIALS 18260445 AND A1820136. (Refer to figure 7-3, sheet 1.)
1. Remove follow-up control (1) from switch
mounting arm (15).
2. Remove flap operating switches (12 and 13)
from switch mounting arm (15). DO NOT disconnect
electrical wiring at switches.
3. Remove knob from control lever (11).
4. Remove remaining items by removing bolt
(18). Use care not to drop parts into tunnel area.
5. Reverse the preceding steps for reinstallation.
Do not overtighten bolt (18) causing lever (11) to bind.
Rig system in accordance with paragraphs 7-18 and
7-19.

Motor Assembly

3. Bolt
4. Transmission Assembly
5. Bolt
6. Nut and Ball Assembly
7. Setscrew
8. Actuating Tube
9. Bolt
10. Bolt
11. Cable Lock
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.

7 8

18
REFER TO FIGURE 7-3
FOR FOLLOW-UP SYSTEM

Attach Bracket
Bolt
Follow-Up Control Bellcrank
Bolt
Drive Pulley
Bolt
Down-Limit Switch
Up-Limit Switch
Electrical Wiring
Snubber Assembly
Bracket
Spacer
Shim
Screw
Setscrew
Switch Adjusting Block
Up-Limit Switch
Switch Actuating Collar
Switch Support
Down-Limit Switch

Detail

NOTES
Use Loctite Sealant, Grade "C" on threads of setscrew (7) after final adjustment.
Ensure shortest end of hinge (2) is at top.
*Beginning with aircraft serials 18259992 and
A182-0117.

AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK150-37

Figure 7-2.

Flap Motor and Transmission Assembly (Sheet 1 of 3)
7-5

24
AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN MODIFIED IN ACCORDANCE WITH SK150-37

~~~~~~~~~~7~~~UP

BEGINNING WITH AIRCRAFT SERIALS
18260699 AND A182-0137

Figure 7-2.
7-6

Flap Motor and Transmission Assembly (Sheet 2 of 3)

position.

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

VIEW

A-A

THIS FLAP ACTUATOR INSTALLATION IS EFFECTIVE
FOR AIRCRAFT SERIALS 18259306 THRU 18260698 AND
A182-0117 THRU A182-0136 WHEN USED AS A REPLACEMENT SPARE FOR SK150-37 OR PRODUCTION FLAP
ACTUATOR INSTALLATIONS PRIOR TO AIRCRAFT
SERIALS 18260699 AND A182-0137

Figure 7-2.

Flap Motor and Transmission Assembly (Sheet 3 of 3)
7-7

b. BEGINNING WITH AIRCRAFT SERIALS 18260446 AND A182-0137. (Refer to figure 7-3, sheet 2.)
1. Remove follow-up control torque tube (32)
from switch mounting arm (15).
2. Remove flap operating switches (12 and 13)
from switch mounting arm (15). DO NOT disconnect
electrical wiring at switches.
3. Remove knob (34) from control lever (11).
4. Remove remaining items by removing bolt
(18). Use care not to drop parts into tunnel area.
5. Reverse the preceding steps for reinstallation.
Do not overtighten bolt (18) causing lever (11) to bind.
Rig system in accordance with paragraphs 7-18 and
7-19.
7-10.

DRIVE PULLEY.

(Refer to figure 7-2.)

7-11. REMOVAL AND INSTALLATION.
a. Remove access plates adjacent to drive pulley
(17) in right wing.
b.. Unzip or remove headliner as necessary for
access to turnbuckles (index 10, figure 7-1), remove
safety wire and loosen turnbuckles.
c. Remove bolt (16) securing flap push-pull rod (12)
to drive pulley (17) and lower RIGHT flap gently.
d. Remove bolt (18) securing actuating tube (8) to
drive pulley (17) and lower LEFT flap gently. Retain
bushing.
e. Remove cable locks (11) securing control cables
to drive pulley (17). Tag cables for reference on
reinstallation.
f. THRU AIRCRAFT SERIALS 18260445 AND A1820136. Remove bolt (9) attaching follow-up control
bellcrank (15) to drive pulley (17).
g. Remove bolt (10) attaching drive pulley (17) to
wing structure.
h. Using care, remove drive pulley through access
opening, being careful not to drop bushing. Retain
brass washer between drive pulley and wing structure
for use on reinstallation. Tape open ends of drive
pulley after removal to protect bearings.
i. To remove left wing drive pulley, use this same
procedure omitting steps "d" and "f."
j. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraphs 7-18 and
7-19, safety turnbuckles and reinstall all items removed for access.
7-12. REPAIR. Repair is limited to replacement of
bearings. Cracked, bent or excessively worn drive
pulleys must be replaced. Lubricate drive pulley
bearings as outlined in Section 2.
7-13.

FLAPS.

(Refer to figure 7-4.)

7-14. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position.
b. Remove access plates (1) from top leading edge
of flap.
c. Disconnect push-pull rod (6) at flap bracket (7).
d. Remove bolts (5) at each flap track. As flap is
removed from wing, all washers, rollers and bushings will fall free. Retain these for reinstallation.
e. Reverse the preceding steps for reinstallation.
If push-pull rod (6) adjustment is not disturbed, rerigging of system should not be necessary. Check
7-8

flap travel and rig, if necessary, in accordance with
paragraphs 7-18 and 7-19.
7-15. REPAIR. Flap repair may be accomplished
in accordance with instructions outlined in Section 18.
7-16.
7-1.)

CABLES AND PULLEYS.

(Refer to figure

7-17. REMOVAL AND INSTALLATION.
a. Remove access plates, fairings, headliner and
upholstery as necessary for access.
b. Remove safety wire, relieve cable tension, disconnect turnbuckles (10) and carefully lower LEFT
flap.
c. Disconnect cables at drive pulleys, remove cable guards and pulleys as necessary to work cables
free of aircraft.
NOTE
To ease routing of cables, a length of wire
may be attached to the end of cable being
withdrawn from the aircraft. Leave wire
in place, routed through structure; then attach the cable being installed and use wire
to pull cable into position.
d. Reverse the preceding steps for reinstallation.
e. After cables are routed in position, install pulleys and cable guards. Ensure cables are positioned
in pulley grooves before installing guards.
f. Re-rig flap system in accordance with paragraphs
7-18 and 7-19, safety turnbuckles and reinstall all
items removed in step "a."
7-18. RIGGING-FLAPS. (Refer to figure 7-2.)
a. Unzip or remove headliner as necessary for access to turnbuckles (index 10, figure 7-1).
b. Remove safety wire, relieve cable tension, disconnect turnbuckles and carefully lower LEFT flap.
c. Disconnect push-pull rods (12) at drive pulleys
(17) in both wings and lower RIGHT flap gently.
d. Disconnect actuating tube (8) from drive pulley
(17).
NOTE
If control cables are not connected to left
and right drive pulleys, actuating tube (8)
and push-pull rods (12) must be disconnected before installing cables. If drive
pulleys (17) are not installed, attach control
cables before installing drive pulleys in the
wings as illustrated in figure 7-5.
e. The 3/32 inch retract cable connects to the forward side of the right drive pulley and to the aft side
of the left drive pulley. The 1/8 inch direct cable
connects to the aft side of the right drive pulley and
to the forward side of the left drive pulley.
f. Adjust both push-pull rods (12) to 8.83±. 12
inches between centers of rod end bearings and
tighten locknuts on both ends. Connect push-pull
rods to flaps and drive pulleys.

Detail A
4

4.
5.

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

Bracket
Spacer

22.
23.

Flaps UP Operating Switch
Insulator
Switch Mounting Arm
Position Indicator
Bushing
Bolt

31. Turnbuckle011
32. Torque Tube
33. Bracket
34. Knob
35. Support
36. Washer (Teflon)

Figure 7-3.

Washer (Metal)
Nylon Guide

WITH AIRCRAFT
S
BEGINNING
SERIALS 18259397 AND A1820099

Flap Control Lever and Follow-Up Installation (Sheet 1 of 2)
Chang

7-9

REFER TO FIGURE 7-2

BEGINNING WITH AIRCRAFT SERIALS 18260446 AND A182-0137

77-10

Change 3

Detail A

BEGINNING WITH AIRCRAFT SERIALS
18261555 THRU 18261971, 18261973 AND
A1820147 & ON.
NOTES

switches (12 and 13) and switch

mounting arm (15).
.

·Apply Loctite Sealant Grade "C", to
threads of knob (34) on installation.

Detail

· BEGINNING WITH
18260683 AND A182-0137

Figure 7-3.
7-10

Change 3

Flap Control Lever and Follow-Up Installation (Sheet 2 of 2)

NOTE
Temporarily connect cables at turnbuckles
(index 10, figure 7-1) and test flaps by hand
to ensure both flaps extend and retract together. If they will not, the cables are incorrectly attached to the drive pulleys. Ensure that the right drive pulley rotates clockwise, when viewed from below, as the flaps
are extended. Tag cables for reference and
disconnect turnbuckles again.
g. AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN NOT MODIFIED IN ACCORDANCE WITH SK150-37 AND WHEN
NOT MODIFIED IN ACCORDANCE WITH FIGURE
7-2, SHEET 3. Screw actuating tube (8) IN toward
transmission (4) by hand to its shortest length (flaps
full up position). Loosen setscrew (7) securing actuating tube (8) to nut and ball assembly (6), hold nut
and ball assembly so that it will not move, hold
RIGHT flap in the full UP position and adjust actuating tube (8) IN or OUT as necessary to align with
attachment hole in drive pulley (17). Tighten set
screw (7) and secure tube to drive pulley with bolt
(18).
h. AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 WHEN MODIFIED
IN ACCORDANCE WITH SK150-37 AND WHEN NOT
MODIFIED IN ACCORDANCE WITH FIGURE 7-2,
SHEET 3. Operate flap motor until actuating tube
(8) is IN to its shortest length (flaps full up position).
Hold RIGHT flap in the full UP position and check
actuating tube (8) to drive pulley (17) attachment holes
for alignment. Operate flap motor toward the DOWN
position until bolt (18) can be installed freely. Loosen setscrew (7) and rotate nut and ball assembly (6)
IN against transmission (4). Tighten setscrew (7)
and bolt (18).
i. BEGINNING WITH AIRCRAFT SERIALS 18260699 AND A182-0137 AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH FIGURE 7-2, SHEET
3. Screw actuating tube (8) IN toward transmission
(4) by hand to .12±. 05 inches between switch actuating collar (30) and transmission as illustrated in
VIEW A-A. Loosen setscrew (7) securing actuating
tube (8) to switch actuating collar (30), hold actuating
collar to maintain . 12±. 05", hold RIGHT flap in the
full UP position and adjust actuating tube (8) IN or
OUT as necessary to align with attachment hole in
drive pulley (17). Tighten setscrew (7) in accordance
with procedures outlined in the following note and
secure tube to drive pulley with bolt (18).
NOTE
Thru Aircraft Serial 18262541 and beginning
with A182-0136: Tighten setscrew (7). Aircraft Serials 18262542 thru 18262544,
18262546 thru 18263011: Apply grade CV
sealant to setscrew (7) threads and torque
to 45 lb-in. Beginning with Aircraft Serial
18263012: Apply grade CV sealant to setscrew (7) threads and torque to 60 lb-in.
If actuating tube (8) is too long to allow
attachment to drive pulley after completion

of steps "g", "h" and "i", proceed to step
j. Disconnect push-pull rod (12) at drive pulley (17),
then connect actuating tube (8) to drive pulley.
k. Manually hold RIGHT flap in full UP position and
readjust push-pull rod (12) to align with attachment
hole in drive pulley. Connect push-pull rod and tighten locknuts.
NOTE
The right flap and actuator must be correctly
rigged before cables and left flap can be rigged.
1. Mount an inclinometer on trailing edge of RIGHT
flap.
NOTE
An inclinometer for measuring control surface
travel is available from the Cessna Service
Parts Center. Refer to figure 6-4.
m. AIRCRAFT SERIALS 18259306 THRU 18260698
AND A182-0117 THRU A182-0136 AND ALL AIRCRAFT NOT MODIFIED IN ACCORDANCE WITH
FIGURE 7-2, SHEET 3.
1. With RIGHT flap in full UP position, adjust
UP-LIMIT switch (20) to operate and shut-off electrical power to motor at degree of travel specified
in figure 1-1.
2. Run RIGHT flap to DOWN position and adjust
DOWN-LIMIT switch (19) to operate and shut-off electrical power to motor at degree of travel specified in
figure 1-1.
n. BEGINNING WITH AIRCRAFT SERIALS 18260699
AND A182-0137 AND ALL AIRCRAFT MODIFIED IN
ACCORDANCE WITH FIGURE 7-2, SHEET 3.
1. With RIGHT flap in full UP position, loosen
setscrew (27) and slide UP-LIMIT switch (29) adjustment block (28) on support (30) to operate switch and
shut-off electrical power to motor at degree of travel
specified in figure 1-1. Tighten setscrew (27).
2. Run RIGHT flap to DOWN position and adjust
DOWN-LIMIT switch (32) adjustment block (28) on
support (31) to operate switch and shut-off electrical
power to motor at degree of travel specified in figure
1-1. Tighten setscrew (27).
o. Run RIGHT flap to full UP position, manually
hold LEFT flap full UP and connect control cables at
turnbuckles (index 10, figure 7-1). Remove reference
tags previously installed in step "f" as turnbuckles
are connected.
p. With flaps full UP, adjust turnbuckles to obtain
30±10 pounds tension on cables. Adjust retract cable
first.
NOTE
Ensure cables are positioned in pulley grooves
and cable ends are positioned correctly at drive
pulleys before tightening turnbuckles.
q. Disconnect push-pull rod at left drive pulley.
Run motor to extend flaps approximately 20 ° and
Change 2

7-11

NOTE
Bushings (4), rollers (3) and spacers (9) are
first positioned through slots in flap tracks,
then are secured to the flap roller supports (2)
with attaching bolts, washers and nuts. Nylon
plug buttons (11) prevent wing flap from chafing
wing trailing edge.
Position spacers (9) and direction of bolts (5) as
required to provide adequate flap clearance at

Detail

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

Access Plate
Flap Support
Roller Assembly
Bushing
Bolt
Push-Pull Rod
Flap Bracket
Bolt
Spacer
Plug Button
Nylon Plug Button

Figure 7-4.
7-12

Flap Installation

OUTBOARD

rDRIVE PULLEY

FWD

TRANSMISSION
- DRIVE PULLEY

PULLEY

SET SCREW

PULLEYS (CABIN TOP)
TURNBUCKLE

RUB STRIP

PUSH-PULL ROD

ACTUATING TUBE

-TURNBUCKLE

TO LEFT

PUSH-PULL ROD
TO RIGHT

WING FLAP

VIEWED

FROM ABOVE

Figure 7-5.

Flap System Schematic

check tension on each flap cable. If necessary, readjust turnbuckles to maintain 30±10 pounds tension
on each cable and safety turnbuckles.
r. Fully retract right flap. Manually hold left flap
in full up position and readjust push-pull rod to align
with attaching hole in drive pulley. Connect push-pull
rod and tighten locknuts.
s. After completion of steps "a" thru "r", operate
flaps and check for positive shut-off of flap motor
through several cycles. Check for specified flap
travel with inclinometer mounted on each flap sepa-

rately.
NOTE
Since the flap rollers may not bottom in the
flap tracks with flaps fully extended, some
free play may be noticed in this position,
7-19. RIGGING-FLAP CONTROL LEVER AND
FOLLOW-UP.
a. THRU AIRCRAFT SERIALS 18260445 AND A1820136. (Refer to figure 7-3, sheet 1. )
1. Disconnect follow-up control rod end (1) at
switch mounting arm (15).
2. Move control lever (11) to full UP position,
then without moving control lever, move switch
mounting arm (15) until cam (10) is centered between
switches (12 and 13). Adjust follow-up control rod
end to align with the attaching hole in the switch
mounting arm and secure rod end to mounting arm
maintaining this position.
3. Adjust flaps DOWN operating switch (12) in
slotted holes until switch roller just clears cam (10)
and secure. This adjustment should provide flaps
down operation to 10°±2 ° and 20°±2 ° .

WING FLAP

4. Adjust flaps UP operating switch (13) in slotted holes for .062 inch clearance between switch
roller and cam (10) when the flaps DOWN operating
switch has just opened in the 10° and 20° position.
NOTE
Flap travel on UP cycle may deviate a maximum of 4 ° from indicated position.
5.

Turn master switch ON and run flaps through

several cycles, stopping at various mid-range settings and checking that cable tension is within limits.
Retract cable tension may increase to 90 pounds when
flaps are fully retracted.
6. Check all rod ends and clevis ends for sufficient thread engagement, all jam nuts are tight and
reinstall all items removed for access.
7. Flight test aircraft and check that follow-up
control does not cause automatic cycling of flaps. If
cycling occurs, readjust operating switches as necessary per steps 3 and 4.
b. BEGINNING WITH AIRCRAFT SERIALS 18260446 AND A182-0137. (Refer to figure 7-3, sheet
2.)
1. Run flaps to full UP position.
2. Remove upholstery and headliner as necessary.
3. Secure follow-up control cable to retract
cable (19) with union assembly (24). Ensure union
assembly is at end of slot in support (20).
4. Pull all slack from follow-up control cable
and with position indicator (16) in full UP position,
connect turnbuckle (31) to follow-up cable.
5. Connect spring (30) to arm assembly (29).
6.

Make minor cable length adjustments using

turnbuckle (31) to position indicator at 0° flaps.
Change 2

7-13

7. With control lever (11) in full up position, adjust switches (12 and 13) in slotted holes until cam
(10) is centered between switch rollers. Be sure
control lever (11) is in full up position during this
adjustment.
8. Mount an inclinometer on trailing edge of one
flap and set to 0 ° . Turn master switch ON and move
control lever to 10 ° position. If flap travel is more
than 10 ° , adjust flaps DOWN operating switch (12)
away from cam (10) and recycle flaps. If flap travel
is less than 10 ° , adjust flaps DOWN operating switch
(12) closer to cam (10) and recycle flaps.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
9. Adjust flaps UP operating switch (13) in slotted holes for .062 inch clearance between switch rol-

SHOP NOTES:

7-14

ler and cam (10) when the flaps DOWN operating
switch has just opened in the 10 ° and 20 ° position.
NOTE
Flap travel on UP cycle may deviate a maximum of 4 ° from indicated position.
10. Turn master switch ON and run flaps through
several cycles, stopping at various mid-range settings and checking that cable tension is within limits.
Retract cable tension may increase to 90 pounds when
flaps are fully retracted.
11. Check all rod ends and clevis ends for sufficient thread engagement, all jam nuts are tight and
reinstall all items removed for access.
12. Flight test aircraft and check that follow-up
control does not cause automatic cycling of flaps. If
cycling occurs, readjust operating switches as necessary per steps 8 and 9.

SECTION 8
ELEVATOR CONTROL SYSTEM

TABLE OF CONTENTS
ELEVATOR CONTROL SYSTEM ......
Description ..............
Trouble Shooting ............
Control Column ............
Elevators ...............
Removal and Installation .....
Repair ..............

8-1. ELEVATOR CONTROL SYSTEM.
figure 8-1.)

...........
Bellcrank ....
Removal and Installation
Arm Assembly .............
Removal and Installation
Cables and Pulleys ...........
Removal and Installation
Rigging ................

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

(Refer to

......
.....

8-2
8-2
8-6
8-6
8-6
8-6
8-6

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 re-rig 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.

Attach push-pull tube correctly.

Cables disconnected.

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

8-1

8-3.

TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

BINDING OR JUMPY MOTION
FELT IN MOVEMENT OF ELEVATOR SYSTEM.

ELEVATORS FAIL TO ATTAIN
PRESCRIBED TRAVEL.

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

Replace defective parts.

Cables slack.

Adjust to tension specified in
figure 8-1.

Cables not riding correctly on
pulleys.

Route cables correctly over pulleys

Nylon grommet on instrument
panel binding.

Replace grommet.

Defective control column
bearing rollers.

Replace defective rollers.

Defective control column
torque tube bearings.

Replace defective bearings.

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

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

Defective elevator hinges.

Replace defective hinges.

Defective pulleys or cable
guards.

Replace defective parts and
install guards properly.

Stops incorrectly set.

Rig in accordance with paragraph 8-14.

Cables tightened unevenly.

Rig in accordance with paragraph 8-14.

Interference at instrument
panel.

Rig in accordance with paragraph 8-14.

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

ELEVATORS.

REMEDY

(Refer to figure 8-2.)

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

e. Using care, remove elevator.
f. To remove left elevator use same procedure,
omitting step "b".
g. Reverse the preceding steps for reinstallation.
8-7. REPAIR. Repair may be accomplished as outlined in Section 18. Hinge bearings may be replaced
as necessary. If repair has affected static balance,
check and rebalance as required.
8-8.

BELLCRANK.

(Refer to figure 8-3.)

NOTE
If trim system is not moved and actuator
screw is not turned, re-rigging of trim
system should not be necessary after reinstallation of elevator.
c. Remove bolts (13) securing elevator torque tubes
(3) to arm assembly (4).
d. Remove bolts (14) from elevator hinges.
8-2

8-9. REMOVAL AND INSTALLATION.
a. Remove access plate below bellcrank on tailcone.

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

5.
6.
7.

Turnbuckle
UP
Elevator
Cable
DOWN
Elevator
Cable

FIGURE 6-2

this system only.
CABLE
30 LBSTENSION:
± 10 LBS (AT AVERAGE TEMPERREFER TO FIGURE 1-1 FOR TRAVEL.

8-3

NOTE

Refer to Section 9 for
trim tab control system.

Detail

14.
15.

D
Figure 8-2.

8-4

Elevator Installation

Bolt
Hinge Bracket

AIRCRAFT SERIALS

Detail

12
13

Figure 8-3.

9.
10.
11.

Push-Pull Tube
Pivot Bolt
Bushing

13.

DOWN Elevator Cable

Elevator Bellcrank Installation

TO
ELEVATOR
UP CABLE

BELLCRANK

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

Figure 8-4.

TO
ELEVATOR
DOWN CABLE

BELLCRANK
STOPS

ELEVATOR
- PUSH-PULL
TUBE

Elevator Bellcrank Travel Stop Adjustment
8-5

b. Remove safety wire, relieve cable tension at
turnbuckles (2) and disconnect turnbuckle eyes at
bellcrank links (3).
c. Disconnect elevator down-springs (5) at bellcrank (4).
d. Disconnect push-pull tube (9) at bellcrank (4).
e. Remove pivot bolt (10) attaching bellcrank (4)
to brackets (8). Remove bellcrank.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 8-14,
safety turnbuckles and reinstall all items removed

place, routed through structure; then attach the cable being installed and pull cable
into position.
f. After cable is routed in position, install pulleys
and cable guards. Ensure cable is positioned in
pulley groove before installing guards.
g. Re-rig system in accordance with paragraph
8-14, safety turnbuckles and reinstall all items
removed in step "a".
8-14.

RIGGING.

(Refer to figure 8-3.)

for access.
8-10.

ARM ASSEMBLY.

(Refer to figure 8-2.)

8-11. REMOVAL AND INSTALLATION.
a. Remove stinger.
b. Remove bolt (10) securing push-pull tube (11) to
arm assembly (4).
c. Remove bolts (13) attaching elevator torque tubes
(3) to arm assembly (4).
d. Remove pivot bolt (12) securing arm assembly
(4) and slide assembly from between elevator torque
tubes.
e. Reverse the preceding steps for reinstallation
and reinstall all items removed for access.
8-12.
8-1.)

CABLES AND PULLEYS.

8-13.

REMOVAL AND INSTALLATION.

(Refer to figure

CAUTION
Position a support stand under tail tie-down
ring to prevent the tailcone from dropping
while working inside.
Remove seats, upholstery and access plates as

b. Remove safety wire and relieve cable tension at
turnbuckles (5).
c. Disconnect cables at control column arm assemblies (index 18, figure 6-2).
d. Disconnect cables at bellcrank links (index 3,
figure 8-3).
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 aircraft. Leave wire in

SHOP NOTES:

8-6

Change 3

CAUTION
Position a support stand under tail tie-down
ring to prevent the tailcone from dropping
while working inside.
a. Streamline elevators, mount an inclinometer on
one elevator and set to 0°.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
b. Adjust bellcrank stop blocks (7) at brackets (8)
to degree of travel specified in figure 1-1.
NOTE
The bellcrank stop blocks (7) are four-sided
bushings, drilled off-center so they may be
rotated to any one of four positions to attain
correct elevator travel. Each 90-degree rotation of the stop, changes the elevator travel
approximately one degree.

in figure 8-1.
d. Check sponge at control column in both UP and
DOWN positions and if necessary, readjust turnbuckles (2) to prevent the control column from hitting
the instrument panel or firewall.
e. Safety turnbuckles and reinstall all items removed for access.

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

b. Adjust bellcrank stop blocks (7) at brackets (8)
to degree of travel specified in figure 1-1.
NOTE
The bellcrank stop blocks (7) are four-sided
bushings, drilled off-center so they may be
rotated to any one of four positions to attain
correct elevator travel. Each 90-degree ro-

tation of the stop, changes the elevator travel

in figure 8-5 by adjusting turnbuckles (2) equally to
tension specified in figure 8-1.
d. Check sponge at control column in both UP and
DOWN positions and if necessary, readjust turnbuckles (2) to prevent the control column from hitting
the instrument panel or firewall.
e. Safety turnbuckles and reinstall all items removed for access.

WARNING

approximately one degree.
c.

Locate elevators in neutral position as illustrated

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

SHOP NOTES:

8-7/(8-8 blank)

SECTION 9
ELEVATOR TRIM TAB CONTROL SYSTEM

TABLE OF CONTENTS

Page

ELEVATOR TRIM TAB CONTROL SYSTEM.
Description .
.......
......
Trouble Shooting
..
.
....
Trim Tab
.
..........
Removal and Installation ......
Trim Tab Actuator ...........
Removal and Installation ......
Disassembly
............
Cleaning, Inspection and Repair ..
Reassembly ............
Trim Tab Free-Play Inspection . .
Trim Tab Control Wheel ........
9-1.

.
.
.

.
.

9-1
9-1
9-1
9-2
9-2
9-2
9-2
9-2
9-5
9-5
9-6
9-6

Removal and Installation .....
. 9-6
Cables and Pulleys .......
...
. 9-6
Removal and Installation ......
9-6
Pedestal Cover ...
.........
9-6
Removal and Installation ......
9-7
Rigging ...................
9-7
Electric Trim Assist Installation
....
. 9-8
Description ............
9-8
Trouble Shooting ..........
9-8
Removal and Installation ......
9-8
Clutch Adjustment .........
9-8
Rigging - Electric Trim Assist . ..
9-10

ELEVATOR TRIM TAB CONTROL SYSTEM.

trim control wheel by means of roller chains, cables,
an actuator and a push-pull tube. A mechanical
pointer, adjacent to the trim wheel indicates tab
position. A "nose-up" setting results in a tab-down
position. Beginning with serial 18264296, an electric
trim assist system may be installed. This system is
described in paragraph 9-15.

9-2. DESCRIPTION. The elevator trim tab, located on the trailing edge of the right elevator, is
controlled by a trim wheel mounted in the pedestal.
Power to operate the tab is transmitted from the
9-3.

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

TRIM CONTROL WHEEL MOVES
WITH EXCESSIVE RESISTANCE.

PROBABLE CAUSE

REMEDY

Cable tension too high.

Check and adjust tension as
specified in figure 9-1.

Pulleys binding or rubbing.

Open access plates and check
visually. Install cables correctly.

Cables not in place on pulleys.

Open access plates and check
visually. Install cables correctly.

Trim tab hinge binding.

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

Defective trim tab actuator.

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

Rusty chain.

Check visually.

Replace chain.

Damaged sprocket.

Check visually.

Replace sprockets.

Bent sprocket shaft.

Observe motion of sprockets.
Replace bent sprocket shafts.
Change 3

9-1

9-3.

TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

LOST MOTION BETWEEN
CONTROL WHEEL AND
TRIM TAB.

REMEDY

Cable tension too low.

Check and adjust tension as
specified in figure 9-1.

Broken pulley.

Open access plates and check
visually. Replace defective
pulley.

Cable not in place on pulleys.

Open access plates and check
visually. Install cables correctly.

Worn trim tab actuator.

Remove and replace worn actuator.

Actuator attachment loose.

Check actuator for security.
Tighten as necessary.

TRIM INDICATOR FAILS TO
INDICATE CORRECT TRIM
POSITION.

Indicator incorrectly engaged
on wheel track.

Check visually and reset
indicator as necessary.

INCORRECT TRIM TAB
TRAVEL.

Stop blocks loose or incorrectly
adjusted.

Adjust stop blocks on cables.
Refer to figure 9-2.

9-4.

TRIM TAB.

(Refer to figure 9-1, sheet 2.)

a. Disconnect push-pull tube (16) from horn assembly (17).

Position a support stand under the tail tiedown ring to prevent tailcone from dropping
while working inside.

NOTE
If trim system is not moved and actuator
screw is not turned, re-rigging of system
should not be necessary after installation
of tab.
b. Drill out rivets securing trim tab hinge to elevator and remove trim tab.

b. Disconnect push-pull tube (16) at actuator (12).
c. Remove access plate beneath actuator.
d. Remove chain guard (11) and disengage chain
from actuator sprocket (8).
e. Remove screws attaching clamps (13) to bracket
(10) and remove actuator (12) through access opening.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 9-14, safety
turnbuckle and reinstall all items removed for access.

NOTE
After tab has been removed and if hinge
pin is to be removed, it is necessary to
spread the crimped ends of the hinge before driving out pin. When a pin has
been installed, crimp ends of hinge to
prevent pin from working out.
c. Reverse the preceding steps for reinstallation.
Rig system if necessary in accordance with paragraph
9-14.
9-6. TRIM TAB ACTUATOR.
sheet 2.)

(Refer to figure 9-1,

9-7. REMOVAL AND INSTALLATION.
a. Relieve cable tension at turnbuckle (index 10,
figure 9-1, sheet 1).
9-2

Change 1

9-7A. DISASSEMBLY. (Refer to figure 9-2A.)
a. Remove actuator in accordance with paragraph
9-7.
b. Disassemble actuator assembly (1) as illustrated
in Detail A as follows:
1. Remove chain guard (3) if not previously removed in step "e" of paragraph 9-7.
2. Using suitable punch and hammer, remove
roll pins (8) securing sprocket (5) to screw (9) and
remove sprocket from screw.
3. Unscrew threaded rod end (15) and remove
rod end from actuator.
4. Remove roll pins (10) securing bearings
(6 and 14) at the housing ends.
5. Lightly tap screw (9) toward the sprocket
end of housing, remove bearing (6) and collar (7).

REFER TO FIGURE 9-2

/c '^J

Detail

Figure 9-1.

K\A

1.
2

Bulkhead (Station 17.00)
Cable Guard

5.
6.
7.

Bulkhead (Station 110.00)
Actuator
Trim Tab

9.
10.

Detail F

Cable End
Turnbuckle

Shaded pulleys are used
for this system only.

10 TO 15 LBS (AT AVERAGE TEMPERATURE FOR THE AREA.)
REFER TO FIGURE 1-1 FOR TRAVEL.

Elevator Trim Tab Control System (Sheet 1 of 2)
9-3

THRU AIRCRAFT SERIALS
18261328 AND A182-0136

BEGINNING WITH AIRCRAFT SERIALS 18261329 AND A182-0137

Retainer

Pedestal Structure

9.
10.

Pedestal Cover
Support Bracket

14.
15.
16.
17.
18.

Stabilizer Rear Spar
Brace
Push-Pull Tube
Horn Assembly
Trim Tab

* Safety wired beginning with aircraft
serials 18260415 and A182-0117
Figure 9-1.
9-4

Change 1

* BEGINNING WITH AIRCRAFT SERIAL 18261226, A182-0137
Elevator Trim Tab Control System (Sheet 2 of 2)

1. With elevators in neutral,

set trim tab to neutral (streamlined).

Position stop blocks (2 and 3) against cable ends and secure to cable A.

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

Position stop block (4) against stop block (3) and secure to cable B.

5. Raise trim tab to specified degree, place stop block (1) against stop block (2)
and secure to cable B.

Figure 9-2.

Elevator Trim Tab Travel Stop Adjustment

6. Lightly tap screw (9) in the opposite direction from sprocket end, remove bearing (14), O-ring
(13) and collar (7).
7. It is not necessary to remove retaining rings
-(11).
9-7B. CLEANING, INSPECTION AND REPAIR.
(Refer to figure 9-3.)
a. DO NOT remove bearing (16) from threaded rod
end (15) unless replacement of bearing is necessary.
b. Clean all component parts, except bearing (16),
by washing in Stoddard solvent or equivalent. Do not
clean sealed bearing (16).
c. Inspect all component parts for obvious indications of damage such as stripped threads, cracks,
deep nicks and dents.
d. Check bearings (6 and 14), screw (9) and threaded rod end (15) for excessive wear and scoring.
Dimensions of the parts are as follows:
BEARING (6)
INSIDE DIAMETER
0. 370" MIN.
INSIDE DIAMETER
0. 373" MAX.
BEARING (14)
INSIDE DIAMETER
SMALL HOLE
0.248" MIN.
SMALL HOLE
0.253" MAX.
LARGE HOLE
0. 373" MIN.
LARGE HOLE
0. 380" MAX.
THREADED ROD END (15)
OUTSIDE DIAMETER
(SHANK)
SCREW (9)
OUTSIDE DIAMETER

0.242" MIN.
0.246" MAX.
0.367" MIN.
0. 370" MAX.

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

h. DO NOT attempt to repair damaged or worn
parts of the actuator assembly. Discard all defective items and install new parts during reassembly.
9-7C. REASSEMBLY. (Refer to figure 9-3.)
a. Always discard the following items and install
new parts during reassembly.
1. Bearings (6 and 14)
2. Roll Pins (8 and 10)
3. O-Ring (13)
4. Nuts (2).
b. During reassembly, lubricate collars (7), screw
(9) and threaded rod end (15) in accordance with
Section 2.
c. Press sprocket (5) into the end of screw (9),
align roll pin holes and install new roll pins (8).
d. Slip bearing (6) and collar (7) on screw (9) and
slide them down against sprocket (5).
e. Insert screw (9), with assembled parts, into
housing (12) until bearing (6) is flush with the end of
housing.

Change 1

9-5

When inserting screw (9) into housing (12),
locate the sprocket (5) at the end of housing
which is farther away from the groove for
retaining ring (11).
The bearings (6 and 14) are not pre-drilled
and must be drilled on assembly. The roll
pins (10) are 1/16 inch in diameter, therefore, requiring a 1/16 (0.0625) inch drill.
f. With bearing (6) flush with end of housing (12),
carefully drill bearing so the drill will emerge
from the hole on the opposite side of housing (12).
DO NOT ENLARGE HOLES IN HOUSING.
g. Press new roll pins (10) into pin holes.
h. Insert collar (7), new O-ring (13) and bearing
(14) into opposite end of housing (12).
i. Complete steps "f" and "g" for bearing (14).
j. If a new bearing (16) is required, a new bearing
may be pressed into the boss. Be sure force bears
against the outer race of bearing.
k. Screw the threaded rod end (15) into screw (9).
1. Install retaining rings (11), if they were removed.
m. Test actuator assembly by rotating sprocket (5)
with fingers while holding threaded rod end (15). The
threaded rod end should travel in and out smoothly,
with no indication of binding.
n. Reinstall actuator assembly in accordance with
paragraph 9-7.CA
9-7D. TRIM TAB FREE-PLAY INSPECTION.
a. Place elevators and trim tab in the neutral position.
b. Using moderate pressure, move the trim tab
trailing edge up and down by hand to check free-play.
c. A maximum of . 163" (total motion up and doen)
measured at the trim tab trailing edge is permissible.
d. If the trim tab free-play is less than .163", the
system is within prescribed limits.
e. If the trim tab free-play is more than. 163'
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 the 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 9-6 through 9-7C. Recheck trim
tab free-play.
9-8. TRIM TAB CONTROL WHEEL.
ure 9-1, sheet 2.)

(Refer to fig-

9-9. REMOVAL AND INSTALLATION.
a. Relieve cable tension at turnbuckle (index 10,
figure 9-1, sheet 1).

9-6

Change 1

Position a support stand under the tail tiedown ring to prevent tailcone from dropping
while working inside.
b. Remove pedestal cover (9) in accordance with
paragraph 9-13.
c. Remove screws attaching control wheel retainer
(1) to left side of pedestal structure (4).
d. Remove retainer (1) and indicator (3), using
care not to drop control wheel (6).
e. Disengage roller chain (8) from sprocket (7) and
remove control wheel (6).
NOTE
Removal of the sprocket (7) from control
wheel shaft is not recommended except
for replacement of parts.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 9-14, safety
turnbuckle and reinstall all items removed for access.
9-10. CABLES AND PULLEYS.
sheet 1.)

(Refer to figure 9-1,

9-11. REMOVAL AND INSTALLATION.
a. Remove seats, upholstery, pedestal cover and
access plates as necessary.

Position a support stand under the tail tiedown ring to prevent tailcone from dropping
while working inside.
b. Remove travel stop blocks (8) from control
cables.
c. Disconnect control cables at turnbuckles (10)
and at cable ends (9).
d. Remove cable guards and pulleys as necessary
to work cables free of aircraft. Disengage roller
chains from sprockets to ease cable removal.
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 the cable being installed and pull cable
into position.
e. After cable is routed in position, install pulleys
and cable guards. Ensure cable is positioned in pulley groove before installing guards. Ensure roller
chains are positioned correctly over sprockets.
f. Re-rig system in accordance with paragraph 9-14,
safety turnbuckle and reinstall all items removed in
step "a."
9-12. PEDESTAL COVER.
sheet 2.)

(Refer to figure 9-1,

2. Nut
5.
6.
7.
8.
9.

Sprocket
Bearing
Collar
Pin
Screw

\
\

10.

16.

Bearing

NOTE

Used with electric trim assist installation
Figure 9-2A

Elevator Trim Tab Actuator Assembly

9-13. REMOVAL AND INSTALLATION.
a. Remove fuel selector valve handle and placard.
b. Remove mike and remove mike mounting bracket.
c. Remove cowl flap control knob.
d. Disconnect electrical wiring to pedestal lights.
e. Remove screws securing pedestal cover to
structure and remove cover.
f. Reverse the preceding steps for reinstallation.
9-14.

RIGGING.

(Refer to figure 9-1, sheet 1.)
C.AUTIONl

Position a support stand under the 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 (8) on trim tab cables.
c. Disconnect push-pull tube from actuator (6).
d. Check cable tension and readjust turnbuckle (10)
if necessary.
NOTE
If chains and/or cables are being installed,
permit actuator screw to rotate freely as
chains and cables are connected. Adjust
cable tension and safety turnbuckle (10).

e. (Refer to figure 9-1, sheet 2.) Rotate trim control wheel (6) full forward (nose down). Ensure
pointer (3) does not restrict wheel movement. If
necessary, reposition pointer using a thin screwdriver to pry trailing leg of pointer out of groove.
NOTE
Full forward (nose down) position of trim
wheel is where further movement is preing sprockets or pulleys.

f. With elevator and trim tab both in neutral
(streamlined), mount an inclinometer on tab and set
to 0° . Disregard counterweight areas of elevators
when streamlining. These areas are contoured so
they will be approximately 3° down at cruising speed.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
g. Rotate actuator screw in or out as required to
place trim tab up with a maximum of 2 ° overtravel,
with actuator screw connected to push-pull tube (16).
Change 3

9-7

h. Rotate trim wheel to position trim tab up and
down, readjusting actuator screw as required to obtain overtravel in both directions.
i. Position stop blocks and adjust as illustrated in
figure 9-2 to degree of trim tab travel specified in
figure 1-1.
j. Install pedestal cover and adjust trim tab pointer
to the center of the "TAKE-OFF" triangle with the
trim tab set at 0°.

k. Safety turnbuckle and reinstall all items removed
in step "a. "

9-15. ELECTRIC ELEVATOR TRIM INSTALLATION
BEGINNING WITH SERIAL 18264296. (Refer to figure 9-3.)

electric drive assembly and a chain connecting the
drive assembly to an additional sprocket mounted on
the standard elevator trim actuator. The electric
drive assembly includes a motor, sprockets and a
chain driven solenoid type adjustable clutch. The
electric drive assembly chain connects to the FORWARD sprocket of the trim tab actuator while the
manual trim chain connects to the A FT sprocket of
the actuator. When the clutch or the drive assembly
is not energized, the drive assembly "free wheels"
and, therefore, has no effect on manual operation.

9-16. DESCRIPTION. An electric elevator trim
assist system may be installed consisting of 2
switches mounted on the pilot's control wheel, a circuit breaker mounted in the center instrument pedestal, fuselage wiring running aft to the 12 Volt D. C.
9-17.

TROUBLE SHOOTING.
TROUBLE

SYSTEM INOPERATIVE.

TRIM MOTOR OPERATING TRIM TAB FAILS TO MOVE.

PROBABLE CAUSE

Change 3

REMEDY

Circuit breaker out.

Check visually.

Defective circuit breaker.

Check continuity.
breaker.

Replace defective

Defective wiring.

Check continuity.

Repair wiring.

Defective trim switch.

Check continuity.
switch.

Replace defective

Defective trim motor.

Remove and bench test. Replace
defective motor.

Defective clutch solenoid.

Check continuity.
solenoid.

Improperly adjusted clutch
tension.

Check and adjust spanner nuts
for proper tension.

Disconnected or broken
cable.

Operate manual trim wheel.
Connect or replace cable.

Defective actuator.

Check actuator operation.
Replace actuator.

9-18. REMOVAL AND INSTALLATION. (Refer to
figure 9-3. )
a. Remove covers (12) beneath tab actuator assembly (6) and drive assembly.
b. Disconnect electrical connectors (13 and 14) and
relieve tension on drive chain (8) at turnbuckle (9).
c. Remove chain guard (10) from tab actuator.
d. Remove mounting bolts from drive assembly and
tab actuator and remove units from the aircraft.
e. Reverse preceding steps for reinstallation.
Check system rigging in accordance with paragraph
9-20.
f. Reinstall all items removed for access.
9-8

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

Reset breaker.

Replace

9-19. CLUTCH ADJUSTMENT. (Refer to figure
9-3.)
a. Remove access covers (12) below drive assembly.
b. Remove safety wire and relieve drive chain tension at turnbuckle (9).
c. Disconnect electric motor by unplugging electrical connectors (13) leading to motor assembly.
d. Remove mounting bolts from drive assembly.
It is necessary to remove unit from aircraft to make
necessary adjustments to clutch.

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

7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.

Trim Tab
Push-Pull Tube
Brace
Stabilizer Rear Spar
Mounting Bracket
Tab Actuator Assembly

Detail

Clamp
Drive Chain
Turnbuckle
Chain Guard
Support
Cover
Connector
Connector
Switch - Disengage
Switch - Pitch Trim
Control Wheel
Circuit Breaker

Detail B

Figure 9-3.

Electric Trim Installation (Sheet 1 of 2)
Change 3

9-9

19.

Screw

23.
24.

Nut
Washer

26.

Washer Assembly

28.

Sprocket

32.
33.

Housing
Cover

35.

Mounting Plate

37.

38.
39.
40.

Chain
Bushing
Sprocket

NOTE

Figure 9-3.

Step 3 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 (20) and (21) to
housing (32) and slide the cover down over 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 ON position.
g. Place disengage switch (15) in ON position.
h. Operate pitch trim switch (16) UP or DOWN to
energize the solenoid clutch (41).
i. Attach a spring scale to drive chain and slowly
pull scale till clutch slippage occurs.
NOTE
During step i, attach scale to drive chain
so that sprocket rotates clockwise as viewed
from the drive end to ensure proper clutch
adjustment.
j. Repeat steps h and i several times to break initial friction of clutch.

Change 3

Electric Trim Installation (Sheet 2 of 2)

NOTE

9-10

Detail

k. Repeat step i verly slowly while watching indicator on spring scale. Slippage should occur between
29. 1 and 32. 9 pounds.
L If tension is not within tolerance, loosen OUTSIDE spanner nut (23) which acts as a lock.
m. Tighten INSIDE spanner nut to increase clutch
tension and loosen nut to decrease clutch tension.
n. When clutch tension is within tolerance, tighten
outside spanner nut against inside nut.
o. Connect electrical wiring removed in step 3,
and reinstall drive assembly in aircraft.
p. Rerig trim system in accordance with paragraph
9-20 and reinstall all items removed for access.
9-20. RIGGING - ELECTRIC TRIM ASSIST. (Refer
to figure 9-3. )
a. The standard manual elevator trim system MUST
be rigged in accordance with paragraph 9-14 before
rigging electric trim assist.
b. Move elevator trim tab to full "NOSE UP" position.
c. Locate NAS228 terminal of turnbuckle (9) at a
point 0. 75 inch from drive assembly housing.
d. Adjust AN155 barrel until chain deflection between sprockets is approximately 0. 25 inch.
i. Resafety turnbuckle and reinstall all items removed for access.

SECTION 10
RUDDER AND RUDDER TRIM CONTROL SYSTEMS

TABLE OF CONTENTS

Page

RUDDER CONTROL SYSTEM ........
Description ..............
Trouble Shooting ............
Rudder Pedal Assembly .........
Removal and Installation ......
Rudder ............
Removal and Installation ......
Repair ..............

10-1
10-1
10-1
10-2
10-2
10-2
10-2
10-5

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

Cables and Pulleys ...........
Removal and Installation ......
Rigging ................
RUDDER TRIM AND NOSE WHEEL
STEERING SYSTEM ...........
Description ..............
Trouble Shooting ............
Rigging ................

(Refer to fig-

10-5
10-5
10-6
10-7
10-7
10-7
10-9

prised of the rudder pedals installation, cables and
pulleys, all of which link the pedals to the rudder and
nose wheel steering.

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 re-rig system, refer to paragraph 10-11.
TROUBLE

RUDDER DOES NOT RESPOND
TO PEDAL MOVEMENT.

PROBABLE CAUSE
Broken or disconnected cables.

REMEDY
Open access plates and check
visually. Connect or replace
cables.

10-1

10-3.

TROUBLE SHOOTING (Cont).
TROUBLE

BINDING OR JUMPY MOVEMENT OF RUDDER PEDALS.

PROBABLE CAUSE

REMEDY

Cables too tight.

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

Cables not riding properly on
pulleys.

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

Binding, broken or defective
pulleys or cable guards.

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

Pedal bars need lubrication.

Refer to Section 2.

Defective rudder bar bearings.

If lubrication fails to eliminate
binding. Replace bearing blocks.

Defective rudder hinge bushings.

Check visually.
bushings.

Clevis bolts too tight.

Check and readjust bolts to
eliminate binding.

Steering rods improperly
adjusted.

Rig system in accordance with
paragraph 10-11.

LOST MOTION BETWEEN
RUDDER PEDALS AND
RUDDER.

Insufficient cable tension.

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

INCORRECT RUDDER TRAVEL.

Incorrect rigging.

Rig in accordance with paragraph
10-11.

10-4. RUDDER PEDAL ASSEMBLY.
ure 10-2.)

(Refer to fig-

10-5. REMOVAL AND INSTALLATION.
a. Remove carpeting, shields and soundproofing
from the rudder pedal and tunnel areas as necessary
for access.
b. Disconnect brake master cylinders (15) and
parking brake cables at pilot's rudder pedals.
c. Remove rudder pedals (2) and brake links (5).
d. Remove fairing from either side of vertical fin,
remove safety wire and relieve cable tension by loosening turnbuckles (index 10, figure 10-1).
e. Disconnect cables (6 and 7) from rudder bar
arms (8).
f. Disconnect wiffletree push-pull rods (index 12,
figure 10-5) at rudder bar arms (11).
g. Remove bolts securing bearing blocks (10) and
carefully work rudder bars out of tunnel area.

10-2

Replace defective

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.
h. 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 10-3.)

10-7. REMOVAL AND INSTALLATION.
a. Disconnect tail navigation light.
b. Remove stinger.
c. Remove fairing from either side of vertical fin,
remove safety wire and relieve cable tension by loosening turnbuckles (index 10, figure 10-1. )

1 Detail B
Detail A

18261528, A1820146

B
-FIGURE 10-2

FIGURE 10-5

4NI

torque tube in UP position. DO
NOT cut pin too short.
1.
2.
3.
4.
5.
6.

Cable Guard
Bracket
Spacer
Pulley
Right Aft Cable
Cotter Pin

Detail

MAINTAIN PROPER CONTROL
CABLE TENSION.

7. Left Aft Cable
8. Travel Stop
9. Bellcrank Assembly
10. Turnbuckle
11. Bulkhead (Station 209.00)
12. Bulkhead (Station 110. 00)
Figure 10-1.

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

Rudder Control System
Change 1

10-3

NOTE
Brake links (5), bellcranks (17), brake torque
tubes (14) and attaching parts are not required
unless dual controls ARE installed. When dual
controls ARE NOT installed, hubs (18) are attached to each end of forward and aft rudder bars.

Detail A

HOLE

CLEARANCE HOLE AFT

1.
2.

Anti-Rattle Spring
Pedal

3. Shaft
4.
5.
6.
7.
8.
9.

Spacer
Brake Link
Right Forward Cable
Left Forward Cable
Rudder Bar Arm (For rudder
cable attachment)
Aft Rudder Bar

10. Bearing Block
11.
12.
13.
14.
15.
16.

Rudder Bar Arm (For wiffletree
push-pull rod attachment)
Forward Rudder Bar
Bracket
Brake Torque Tube
Master Cylinder
Bearing

Detail B

17. Bellcrank
18.

Single Controls Hub

Figure 10-2.
10-4

Rudder Pedals Installation

1.2. Upper
Bolt Hinge
4.
5.

Detail

Nut
Center Hinge

B
9

DetailC
BEGINNING WITH AIRCRAFT SERIAL 18261529,
A1820147

9
4
Detail C

Figure 10-3. Rudder Installation

d. Disconnect cables (index 5 and 7, figure 10-1)
from rudder bellcrank.

e. With rudder supported, remove all hinge bolts,
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.

(Refer to figure

10-1.)

10-10. REMOVAL AND INSTALLATION.
a. Remove seats, upholstery and access plates as
necessary.
b. Relieve cable tension at turnbuckles (10) and disconnect cables.
c. Disconnect cables (index 6 and 7, figure 10-2)
from rudder bar arms.
d. Remove cable guards and pulleys as necessary
to work cables free of aircraft.

Change 1

10-5

BLOCK

BLOCK RUDDER HALF

WIRE POINTER

TWEEN STRAIGHTEDGES
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.

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.

Using soft lead pencil, mark rudder at point corresponding to soft wire indexing point (neutral).

Remove straightedges and blocks.

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 10-4.

Checking Rudder Travel

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
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. (Refer to figure 10-5.)
a. Adjust travel stop bolts (index 8, figure 10-1) to
attain correct rudder travel as specified in figure 1-1.
Figure 10-4 illustrates correct travel and one method
of checking.

10-6

b. THRU AIRCRAFT SERIALS 18261328 AND A1820136. Remove rudder trim chain (10) by removing
the lower screws from support bracket (7), using
care not to drop washers (20). These washers are
used as shims to adjust chain (10) tension by raising
or lowering support bracket (7). Spring bracket (7)
downward until chain (10) can be disengaged from
sprockets (9 and 19).
c. BEGINNING WITH AIRCRAFT SERIALS 18261329 AND A182-0137. Loosen adjustable idler
sprocket (25) and disengage chain from sprockets
(9 and 19).
d. Disconnect steering bungee adjustable rod end
(26) from wiffletree (14).
e. Remove fairing from either side of vertical fin,
remove safety wire and relieve cable tension at turnbuckles (index 10, figure 10-1).
f. Clamp rudder pedals in neutral position and center wiffletree (14) by adjusting push-pull rods (12).
Wiffletree is centered when the bolts in each end are
the same distance from the bulkhead just forward of
the wiffletree. Tighten jam nuts.

g. Maintaining rudder pedals in neutral position,
adjust turnbuckles (index 10, figure 10-1) to specified tension with the rudder offset one degree to the
right, (5/16 inch at lower trailing edge). Safety
turnbuckles.

10-13. DESCRIPTION. A sprocket-operated screw
mechanism to provide rudder trim is incorporated at
the aft end of the steering bungee (16). The trim
system is operated by a trim control wheel (4),
mounted in the pedestal. Nosewheel steering is
accomplished through use of the rudder pedals. The
steering bungee (16) links the nose gear to a wiffletree (14) which is operated by push-pull rods (12)
connected to the rudder pedal bar arms (13).

NOTE
After completing the preceding steps, the rudder control system is rigged. The rudder control system MUST be correctly rigged prior to
rigging the rudder trim and nosewheel steering
system. Refer to paragraph 10-15 for rigging
the rudder trim and nosewheel steering system.

NOTE
The rudder control system, rudder trim control system and nosewheel steering systems
are interconnected. Adjustments to any one
of these systems will affect the others. For
maintenance to the nose gear steering, other
than rigging, refer to Section 5.

10-12. RUDDER TRIM AND NOSEWHEEL STEERING SYSTEM. (Refer to figure 10-5.)

10-14.

TROUBLE SHOOTING.
NOTE
This trouble shooting chart should be used in
conjunction with the trouble shooting chart in
paragraph 10-3.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 10-15.
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 paragraph 10-15.

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 paragraph 10-15.

10-7

THRU AIRCRAFT
SERIALS
18261328
AND A182-0136

12.
13.
14.

Push-Pull Rod
Rudder Bar Arm
Wiffletree (Bellcrank)

22.B. Pedestal
Structure
Trim Shaft
Bearing
pwper Bearing
24
Lowe Berig
24

r The free play of chain (10) at midpoint (neutral position) should be
approximately 1/2 inch thru air-

Figure 10-5.
10-8

Rudder Trim Control System

BEGINNING WITH AIRCRAFT

10-15.

RIGGING.

(Refer to figure 10-5.)
NOTE

The rudder control system MUST be correctly
rigged prior to rigging the rudder trim and
nosewheel steering system. Refer to paragraph 10-11 for rigging the rudder control system.
a. After completing step "g" of paragraph 10-11,
tie down or weight tail to raise nosewheel free of
ground.
b. Extend strut and ensure nose gear is centered
against external centering stop. (Refer to note in
figure 5-7. )
c. With rudder pedals clamped in neutral position,
adjust steering bungee rod end (26) to .81 + .00 -. 06
inch from the aft face of sprocket (19). Maintaining
this adjustment, rotate sprocket (19) IN or OUT as
necessary to align rod end with attaching hole in wif-

fletree (14) and install.
d. Rotate trim control wheel (4) until indicator (2)
is centered in pedestal slot (neutral).
e. Without moving sprocket (19), engage chain on
sprockets (9 and 19).
f. THRU AIRCRAFT SERIALS 18261328 AND A1820136. Tighten chain to approximately 1/2 inch free
play at its mid-point by adding washers (20) as required, then install lower screws in bracket (7).
g. BEGINNING WITH AIRCRAFT SERIALS 18261329 AND A182-0137. Tighten chain by adjusting
idler sprocket (25).
h. Lower nosewheel to ground, remove clamps
from rudder pedals, tighten all jam nuts and reinstall
all items removed for access.

WARNING
Be sure rudder moves in the correct direction when operated by the rudder pedals and
trim control wheel.

SHOP NOTES:

10-9/(10-10 blank)

SECTION 11
ENGINE

TABLE OF CONTENTS
ENGINE COWLING .............
Description ..............
Removal and Installation ........
Cleaning and Inspection .......
Repair ................
....
..
....
Cowl Flaps ..
Description ............
Removal and Installation ......
Rigging ..............
ENGINE ..................
Description ..............
Engine Data ..............
Trouble Shooting ............
Static Run-Up Procedures . ......

Page
11-2
11-2
11-2
. . 11-2
11-2
11-2
..
11-2
11-2
11-2
11-2
11-2
11-3
11-6
.11-8A

.11-10
Cleaning and Inspection ......
.
Removal and Installation . . . . .11-11
11-11
Engine Mount .............
11-11
Description ............
.11-11
Removal and Installation . . . .
.11-11
Repair .............
.11-11
Engine Shock-Mount Pads ......
11-11
ENGINE OIL SYSTEM ..........
11-11
.....
..
Description
11-13
Trouble Shooting ..........
11-15
Full-Flow Oil Filter ........
11-15
Description ..........
. . . .11-15
Removal and Installation
.11-17
Filter Adaptor ..........
.11-17
..........
Removal
Disassembly, Inspection and
.11-17
........
Reassembly
11-18
Installation ..........
.11-18
Oil Cooler ...........
11-18
Description ..........
.11-18
.......
ENGINE FUEL SYSTEM
11-18
Description ............
11- 18
Carburetor ............

Removal and Installation .
Idle Speed and Mixture
.........
Adjustments
INDUCTION AIR SYSTEM ........
.
...........
Description
. ..
..
Airbox . . ....
Removal and Installation .
Cleaning and Inspection . .
Induction Air Filter .........
.........
.
Description
Removal and Installation. .
Cleaning and Inspection .....
.........
.
IGNITION SYSTEM
.........
Description .

Description
Rigging

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

.. 11-18

11-18
11-19
11-19
. .. 11-19
. .. 11-19
. .11-19
.
11-19
11-19
. ..11-19
11-19
11-19
.11-19

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

11-25
1-25

11-26
Mixture Control ........
Carburetor Heat Control . .. . 11-26
11-26
......
Propeller Control
11-26
STARTING SYSTEM ..........
.11-27
Description ...........
11-27
Trouble Shooting ..........
11-28
Primary Maintenance ........
11-28
Starter Motor ...........
.
Removal and Installation . . .11-28
. .11-28
EXHAUST SYSTEM .........
11-28
..
Description ..........
11-28
Removal and Installation ......
.. 11-28
...........
Inspection.
EXTREME WEATHER MAINTENANCE . .11-30
11-30
Cold Weather ...........
11-30
Hot Weather ............
11-30
Seacoast and Humid Areas ......
.11-30
Dusty Areas ...........
11-30
......
Ground Service Receptacle
11-30
............
Hand-Cranking.

Change 2

11-1

11-1.

ENGINE COWLING.

11-2. DESCRIPTION. The engine cowling is divided
into two major removable segments. The 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
and remote strainer drain control. Controllable cowl
flaps are attached to the trailing edge of the lower
cowl segment to aid in controlling the engine temperature. Screws fasten the upper and lower segments
together at the nose cap. Quick-release fasteners
are used along the parting surfaces and aft end, allowing the removal of either segment individually. Beginning with aircraft serial 18260826, cowl-mounted
landing and taxi lights are mounted in the lower cowling nose cap. Beginning with aircraft serial 18261426, instead of attaching directly to the fuselage,
the cowling attaches to shock-mounts, which in turn,
are fastened to the fuselage.
11-3. REMOVAL AND INSTALLATION.
a. Disconnect cowl flap control devises at cowl
flaps.
b. Remove screws securing upper and lower cowling segments together at the nose cap.
c. Release the quick-release fasteners attaching
the cowling to the fuselage and at the parting surfaces
of the upper and lower segments.
d. (BEGINNING WITH AIRCRAFT SERIAL 18260826.) Disconnect the landing and taxi light wires
at the quick-disconnects.
e. Disconnect air induction duct on lower cowl segment at airbox and carefully remove cowling.
f. 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. Check
cowl flap rigging and re-rig, if necessary, in accordance with paragraph 11-9.
11-4. 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.
11-5. 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.
11-2

11-6.

COWL FLAPS.

11-7. 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 aft edge of the lower cowl segment.
11-8. REMOVAL AND INSTALLATION. (Refer to
figure 11-1.)
a. Place cowl flap control lever (11) in the OPEN
position.
b. Disconnect cowl flap control devises (6) from
cowl flap shock-mounts (7).
c. Remove safety wire securing hinge pins 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 11-9.
11-9. RIGGING. (Refer to figure 11-1.)
a. Disconnect cowl flap control devises (6) from
cowl flap shock-mounts (7).
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 cowl flap control lever (11) in the CLOSED
position. If the control lever cannot be placed in the
closed position, loosen clamp (3) at upper end of controls and slip housings in clamp or adjust controls at
upper clevis (10) to position control lever in bottom
hole of position bracket (9).
d. With the control lever in CLOSED position, hold
one cowl flap closed, streamlined with trailing edge
of lower cowl. Loosen jam nut and adjust clevis (6)
on the control to hold cowl flap in this position and
install bolt.
NOTE
If the lower control clevis (6) cannot be adjusted far enough to streamline flap and
still maintain sufficient thread engagement,
loosen the lower control housing clamp (4)
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.
f. When the cowl flaps are lowered, they should be
open 13°+3°-1 ° measured in a straight line from the
fuselage to the trailing edge of cowl flaps.
g. Check that all clamps and jam nuts are tight.
11-10.

ENGINE.

11-11. DESCRIPTION. An air cooled, wet-sump,
six-cylinder, horizontally-opposed, direct-drive,
carbureted, Continental 0-470 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 paragraph 11-12 for engine data. For repair
11-12.

and overhaul of the engine, accessories and propeller,
refer to the appropriate publications issued by their
manufacturer's. These publications are available
from the Cessna Service Parts Center.

ENGINE DATA.

Aircraft Series

182 and SKYLANE

MODEL (Continental)

O-470-R

Rated Horsepower at RPM

230 at 2600

Number of Cylinders

6 Horizontally-Opposed

Displacement
Bore
Stroke

470 Cubic Inches
5.00 Inches
4.00 Inches

Compression Ratio

7.00:1

Magnetos
Right Magneto
Left Magneto

Slick No. 662
Fires 22 ° BTC, Lower Left, Upper Right
Fires 22 ° BTC, Upper Left, Lower Right

Firing Order

1-6-3-2-5-4

Spark Plugs

18 MM (Refer to current Continental active
factory approved spark plug chart.)
330 ± 30 LB-IN.

Torque
Carburetor (Marvel-Schebler)

MA-4-5

Tachometer

Mechanical Drive

Oil Sump Capacity
With External Filter

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

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

30-60
10
100
Between No. 2 and No. 4 Cylinder

Oil Temperature
Normal Operating
Maximum
Probe location

Within Green Arc
Red Line (225°F.)
Below Oil Cooler

Cylinder Head Temperature
Normal Operating
Maximum
Probe Location

Within Green Arc
Red Line (460°F.)
Lower side of Number 6 Cylinder (THRU SERIALS
18260055 AND A182-0116.
Lower side of Number 3 Cylinder (18260056 THRU
18260825 AND A182-0117 AND ON.)
Lower side of Number 2 Cylinder (18260826
THRU 18262465.)
Lower side of Number 1 Cylinder (BEGINNING
WITH 18262466.

Direction of Crankshaft
Rotation (Viewed from Rear)

Clockwise

Dry Weight-With Accessories

438 LB (Weight is approximate and will vary
with optional accessories installed.)
Change 2

11-3

11-12. ENGINE DATA. (Cont.)
Aircraft Series

182 and Skylane

MODEL (Continental)

0-470-S

Rated Horsepower at RPM

230 at 2600

Number of Cylinders

6-Horizontally-Opposed

Displacement
Bore
Stroke

470 Cubic Inches
5. 00 Inches
4. 00 Inches

Compression Ratio

7. 00:1

Magnetos
Right Magneto

Slick No. 662
Fires 22 ° BTC, Lower Left,
Upper Right
Fires 22 ° BTC, Upper Left
Lower Right

Left Magneto

Firing Order

1-6-3-2-5-4

Spark Plugs

18 MM (Refer to current
Continental active factory
approved spark plug chart.)
330 # 30 LB-IN.

Torque
Carburetor (Marvel-Schebler)

MA-4-5

Tachometer

Mechanical Drive

Oil Sump Capacity
With External Filter

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

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

30-60
10
100
Between No. 2 and No. 4 Cyl.

Oil Temperature
Normal Operating
Maximum
Probe Location

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

Cylinder Head Temperature
Normal Operating
Maximum
Probe Location

Within Green Arc
Red Line 460°F.)
Lower side of Number 3 Cylinder

Direction of Crankshaft
Rotation (Viewed from Rear)

Clockwise

Dry Weight-With Accessories

438 LB (Weight is approximate and
will vary with optional accessories
installed.)

NOTE
The 0-470-S engine is an acceptable replacement for the 0-470-R engine beginning with aircraft
serial 18260826. When the 0-470-S engine is installed, SK182-50 must be complied with.
11-4

Change 2

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

Pedestal
Cowl Flap Control
Clamp
Clamp
Cowl Flaps
Clevis
Shock-Mount
Bracket
Position Bracket
Clevis
Control Lever
Bushing

Refer to section 2 for
Cowl-Flap hinge ins\pection, (Also refer tow
Detail
C
Detail C

service letter SE 71-27)
Figure 11-1.

Detail
B
Detail D

Cowl Flaps Installation
Change 2

11-5

11-13.

TROUBLE SHOOTING.
TROUBLE

ENGINE WILL NOT START.

11-6

Change 2

PROBABLE CAUSE

REMEDY

Improper use of starting procedure.

Review starting procedure.
to Owner's Manual.

Fuel tanks empty.

Visually inspect cells. Fill
with proper grade and quantity
of gasoline.

Mixture control in the IDLE
CUT-OFF position.

Move control to the full RICH
position.

Fuel selector valve in OFF
position.

Place selector valve in the ON
position to a cell known to
contain gasoline.

Defective carburetor.

If engine will start when primed
but stops when priming is discontinued, with mixture control
in full RICH position, the carburetor is defective. Repair or
replace carburetor.

Carburetor screen or fuel
strainer plugged.

Remove carburetor and clean
thoroughly. Refer to paragraph
11-48.

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

Refer to paragraph 11-89.

Engine flooded.

Refer to paragraph 11-89.

Water in fuel system.

Open fuel strainer drain and
check for water. If water is
present, drain fuel cell sumps,
lines, strainer and carburetor.

Defective aircraft fuel system.

Refer to Section 12.

Fuel contamination.

Drain all fuel and flush
out fuel system. Clean all
screens, fuel lines, strainer
and carburetor.

Defective ignition system.

Refer to paragraph 11-67.

Defective magneto switch or
grounded magneto leads.

Check continuity. Repair or
replace switch or leads.

Spark plugs fouled.

Remove, clean and regap plugs.
Test harness cables to persistently
fouled plugs. Replace if defective.

Refer

11-13.

TROUBLE SHOOTING (Cont).
TROUBLE

ENGINE STARTS BUT
DIES, OR WILL NOT
IDLE.

ENGINE RUNS ROUGHLY,
WILL NOT ACCELERATE
PROPERLY, OR LACKS
POWER.

PROBABLE CAUSE

REMEDY

Idle stop screw or idle mixture
incorrectly adjusted.

Refer to paragraph 11-49.

Carburetor idling jet plugged.

Clean carburetor and fuel strainer.
Refer to paragraph 11-48.

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 cell sumps, lines,
strainer and carburetor.

Defective ignition system.

Refer to paragraph 11-67.

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

Refer to paragraph 11-89.

Induction air leaks.

Check visually.
cause of leaks.

Manual primer leaking.

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

Leaking float valve or float
level set too high.

Perform an idle mixture check.
Attempt to remove any rich
indication with the idle mixture
adjustment. If the rich indication cannot be removed, the
float valve is leaking or the
float level is set too high. Replace defective parts, reset
float level.

Defective carburetor.

If engine will start when primed
but stops when priming is discontinued, with mixture control
in full RICH position, the carburetor is defective. Repair or
replace carburetor.

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 fuel system.

Refer to Section 12.

Restriction in aircraft fuel
system.

Refer to Section 12.

Worn or improperly rigged
throttle or mixture control.

Check visually. Replace worn
linkage. Rig properly.

Correct the

Change 2

11-7

11-13.

TROUBLE SHOOTING (Cont).

TROUBLE
ENGINE RUNS ROUGHLY,
WILL NOT ACCELERATE
PROPERLY, OR LACKS
POWER. (Cont.)

POOR IDLE CUT-OFF.

11-8

Change 2

PROBABLE CAUSE

REMEDY

Spark plugs fouled or improperly gapped.

Remove, clean and regap plugs.
Replace if defective.

Defective ignition system.

Refer to paragraph 11-67.

Defective or badly adjusted
accelerating pump in carburetor.

Check setting of accelerating
pump linkage and adjust as
necessary.

Float level set too low.

Check and reset float level.

Defective carburetor.

If engine will start when primed
but stops when priming is discontinued, with mixture control
in full RICH position, the carburetor is defective. Repair or
replace carburetor.

Defective engine.

Check compression. Listen for
unusual engine noises. Engine
repair is required.

Restricted carburetor air
filter.

Check visually. Clean in
accordance with Section 2.

Cracked engine mount.

Inspect and repair or replace
mount as required.

Defective mounting bushings.

Inspect and install new bushings
as required.

Propeller control in high
pitch (low rpm) position.

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

Fuel contamination.

Check all screens in fuel system.
Drain all fuel and flush out system. Clean all screens, lines,
strainer and carburetor.

Worn or improperly rigged
mixture control.

Check that idle cut-off stop on
carburetor is contacted.
Replace worn linkage. Rig
properly.

Manual primer leaking.

Disconnect primer outlet line.
If fuel leaks through primer,
it is defective. Repair or
replace primer.

Defective carburetor.

Repair or replace carburetor.

11-13A. STATIC RUN-UP PROCEDURES. In a
case of suspected low engine power, a static RPM
run-up should be conducted as follows:
a. Run-up engine, using take-off power and 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.

3. Check magneto timing, spark plugs and
ignition harness for settings and conditions.
4. On fuel injection engines, check fuel injection
nozzles for restriction and check for correct unmetered fuel flow.
5. Check condition of induction air filter. Clean
if required.
6. Perform an engine compression check (Refer
to engine Manufacturer's Manual).

NOTE
Daily changes in 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 2575 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 13 for procedures. )
NOTE
If verification of governor operation is
necessary, the governor may be removed
from the engine and a flat plate installed
over the engine pad. Run-up engine to
determine that governor was adjusted
properly.
2. Check carburetor heat control (carburetor
equipped engines) for proper rigging. If partially
open it would cause a slight power loss. On fuel injected engines check operation of alternate air door
spring or magnetic lock to make sure door will remain closed in normal operation.

11-14. 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 with the engine mount attached to 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.
b. Place fuel selector valve in the OFF position.
c. Remove engine cowling in accordance with paragraph 11-3.
d. Disconnect battery cables and insulate terminals
as a safety precaution.
e. Drain fuel strainer and lines with strainer drain
control.

SHOP NOTES:

Change 2

11-8A

NOTE
During the following procedures, remove
any clamps or lacings which secure controls, wires, hoses or lines to the engine,
engine mount 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 13. Cover exposed end of crankshaft
flange and propeller flange to prevent entry of foreign
material.
i. Disconnect throttle and mixture controls at carburetor. 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 propeller governor control at governor. Note EXACT position, size and number of attaching washers for reference on reinstallation. Remove clamps attaching control to engine and pull control aft clear of engine.
k. Disconnect all hot and cold air flexible ducts
and remove.
1. Remove exhaust system in accordance with paragraph 11-85.
m. Disconnect carburetor heat control from arm on
airbox. Remove clamps and pull control clear of
engine.

SHOP NOTES:

11-8B

Change 2

n. Disconnect wires and cables as follows:
1. Disconnect tachometer drive shaft at adapter.
CAUTION
When disconnecting starter cable do not
permit starter terminal bolt to rotate.
Rotation of the bolt could break the conductor between bolt and field coils causing the starter to be inoperative.
2. Disconnect starter electrical cable at starter.
3. Disconnect cylinder head temperature wire at
probe.
4. Disconnect carburetor air temperature wires
at quick-disconnects.
5. Disconnect electrical wires and wire shielding ground at alternator.
6. Disconnect exhaust gas temperature wires at
quick-disconnects.
7. Remove all clamps and lacings attaching
wires or cables to engine and pull wires and cables
aft to clear engine.
o. Disconnect lines and hoses as follows:
1. Disconnect vacuum hose at vacuum pump.
2. Disconnect oil breather and vacuum system
oil separator vent lines where secured to the engine.
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 oil temperature bulb below cooler.
4. Disconnect primer line at firewall fitting.
5. Disconnect fuel supply hose at fuel strainer.
6. Disconnect oil pressure line at firewall
fitting.
7. Disconnect manifold pressure line at firewall.
p. 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.

- A U T IO N I
Place a suitable stand under tail tie-down
ring before removing engine. The loss of
engine weight will cause the aircraft to be
tail heavy.
q. 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 mount pads.
r. Remove bolts attaching engine to engine mount
pads and slowly hoist engine and pull it forward.
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.
s. Remove engine shock-mount pads and bonding
straps.
11-15. CLEANING. The engine may be cleaned with
Stoddard solvent or equivalent, then dried thoroughly.

CAUTION
Particular care should be given to electrical
equipment before cleaning. Cleaning fluids
should not be allowed to enter magnetos,
starter, alternator, etc. Protect these components before saturating the engine with solvent. All other openings should also be covered before cleaning the engine assembly.
Caustic cleaning solutions should be used
cautiously and should always be properly
neutralized after their use.
11-16. 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.
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.
11-17. 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
through protective plys, cuts, breaks, stiffness,

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 fittings or
severe discoloration of the hoses.
NOTE
Avoid excessive flexing and sharp bends
when examining hoses for stiffness.
e. All flexible fluid carrying hoses in the engine
compartment should be replaced at engine overhaul
or every five years, whichever occurs first.
f. For major engine repairs, refer to the manufacturer's overhaul and repair manual.
11-18. BUILD-UP. Engine build-up consistsof installation 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.
11-19. INSTALLATION. Before installing the engine
on the aircraft, install any items which were removed
from the engine or aircraft after the engine was removed.
NOTE
Remove all protective covers,
and identification tags as each
nected or installed. Omit any
present on a particular engine

plugs, caps
item is conitems not
installation.

a. Hoist the engine to a point near the engine mount.
b. Install engine shock-mount pads and bonding
straps as illustrated in figure 11-2.
c. Carefully lower engine slowly into place on the
engine mount. Route controls, lines, hoses and
wires in place as the engine is positioned on the engine mount pads.
NOTE
Be sure engine shock-mount pads, 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 450-500 Ib-in.
e. Route throttle, mixture and propeller controls
to their respective units and connect. Secure controls in position with clamps.
f. Route carburetor heat control to airbox and connect. Secure control in position with clamps.
NOTE
Throughout the aircraft fuel system, from the
fuel cells to the carburator, use NS-40 (RAS-4)
(Snap-On-Tools Corp., Kenosha, Wisconsin),
Change 3

11-9

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, 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.
g. Connect lines and hoses as follows:
1. Connect manifold pressure line at firewall
fitting.
2. Connect oil pressure line at firewall fitting.
3. Connect fuel supply hose at fuel strainer.
4. Connect primer line at firewall fitting.
5. Connect oil temperature bulb below cooler.
6. Connect oil breather and vacuum system oil
separator vent lines where secured to the engine.
7. Connect vacuum hose at vacuum pump.
h. Connect wires and cables as follows:
1. Connect electrical wires and wire shielding
ground at alternator.
2. Connect cylinder head temperature wire at

o. Check all switches are in the OFF position and
connect battery cables.
p. Rig engine controls in accordance with paragraphs 11-73, 11-74, 11-75 and 11-76.
q. Inspect engine installation for security, correct
routing of controls, lines, hoses and electrical wiring, proper safetying and tightness of all components.
r. Install engine cowling in accordance with paragraph 11-3. Rig cowl flaps in accordance with paragraph 11-9.
s. Perform an engine run-up and make final adjustments on the engine controls.
11-20.

FLEXIBLE FLUID HOSES.

11-21. LEAK TEST.
a. After each 50 hours of engine operation, all flexible fluid hoses in the engine compartment should be
checked for leaks as follows:
1. Examine the exterior of hoses for evidence of
leakage or wetness.
2. Hoses found leaking should be replaced.
3. Refer to paragraph 11-17 for detailed inspection procedures for flexible hoses.

probe.
11-22.

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.
5. Connect exhaust gas temperature wires and
carburetor air temperature wires at quick-disconnects.
6. Install clamps and lacings securing wires
and cables to engine, engine mount and brackets.
i. Install exhaust system in accordance with paragraph 11-85.
j. Connect all hot and cold air flexible ducts.
k. Install propeller and spinner in accordance wtih
instructions outlined in Section 13.
1. 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.
m. Clean and install induction air filter in accordance with Section 2.
n. 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.

11-10

Change 3

REPLACEMENT.

or loosening of the nut.
b. Provide as large a bend radius as possible.
c. Hoses should have a minimum of one-half inch
clearance 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 AC 43.13-1, Chapter 10, for additional
installation procedures for flexible fluid hose assemblies.
11-23.

ENGINE BAFFLES.

11-24. 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. These 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.
11-25. CLEANING AND INSPECTION. The engine
baffles should be cleaned with a suitable solvent to
remove oil and dirt.

5. Nut
7.

Barrel Nut

8. Roll Pin
9. Spacer
10. Ground Strap

MOUNT-TO-FIREWALL

6
TORQUE MOUNT-TO-FIREWALL
BOLTS TO 160-190 LB-IN
ENGINE-TO-MOUNT
* Washer (3) is installed on the lower
mounts only beginning with aircraft
serials 18260291 and A182-0117.
Figure 11-2.

Engine Mount Details

NOTE
The rubber-asbestos seals are oil and grease
resistant but should not be soaked in solvent
for long periods,
Inspect baffles for cracks in the metal and for loose
and/or torn seals. Repair or replace any defective
parts.
11-26. 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.
11-27. 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.
11-28.

ENGINE MOUNT.

TORQUE ENGINE-TO-MOUNT
BOLTS TO 450-500 LB-IN

b. Remove bolts from upper and lower mount-tofuselage structure and carefully remove engine mount.
c. Reverse the preceding steps for reinstallation.
Torque bolts to 160-190 lb-in. Reinstall engine in
accordance with paragraph 11-19.
11-31. REPAIR. Repair of the engine mount shall
be performed carefully as outlined in Section 18.
The mount shall be painted with heat-resistant black
enamel after welding or whenever the original finish
has been removed. This will prevent corrosion.
11-32. ENGINE SHOCK-MOUNT PADS. (Refer to
figure 11-2.) The bonded rubber and metal shockmounts are designed to reduce transmission of engine vibrations to the airframe. The rubber pads
should be wiped clean with a clean dry cloth.
NOTE
Do not clean the rubber pads and dampener
assembly with any type of cleaning solvent.
Inspect the metal parts for cracks and excessive wear
due to aging and deterioration. Inspect the rubber
pads for separation between the pad and metal backing,
swelling, cracking or a pronounced set of the pad.
Install new parts for all parts that show evidence of
wear or damage.

(Refer to figure 11-2.)

11-29. DESCRIPTION. The engine mount is composed of sections of steel tubing welded together and
reinforced with gussets. The mount is fastened to
the fuselage at four points. The engine is attached
to the engine mount with shock-mount assemblies
which absorb engine vibrations. Each engine mount
pad has a small hole for a locating pin which serves
as a locating dowel for the engine shock-mounts.
11-30. REMOVAL AND INSTALLATION.
a. Remove engine in accordance with paragraph
11-14.

11-33.
11-3.)

ENGINE OIL SYSTEM.

(Refer to figure

11-34. DESCRIPTION. A wet-sump, pressurelubricating oil system is employed in the engine.
Oil under pressure from the oil pump is fed through
drilled crankcase passages which supply oil to the
crankshaft main bearings and camshaft bearings.
Connecting rod bearings are pressure-lubricated
through internal passages in the crankshaft. Valve
mechanisms are lubricated through the hollow push-

11-11

THERMOSTAT

THERMOSTAT
--

XOPEN

STANDARD
OIL COOLER

PLUG

CLOSED

STANDARD
OIL COOLER

THERMOSTAT
OPEN

NON-CONGEALING
OIL COOLER

TO
PROPELLER

PROPELLER
CONTROL

PROPELLER
GOVERNOR

OIL
TEMPERATURE

CAP

CODE:

SUMP OIL, RETURN
OIL, AND SUCTION
OIL

OPTIONAL EXTERNAL
FILTERI
Figure 11-3.

11-12

Engine

Schematic
Oil

rods, which are supplied with oil from the crankcase
oil passages. The propeller is supplied oil, boosted
by the governor through the forward end of the crankshaft. Oil is returned by gravity to the engine oil
sump. Cylinder walls and piston pins are spraylubricated by oil escaping from connecting rod bearings. The engine is equipped with an oil cooler and

11-35.

a thermostat valve to regulate engine oil temperature.
A pressure relief valve is installed to maintain proper oil pressure at higher engine speeds. Removable
oil filter screens are provided within the oil system.
An external, replaceable element oil filter is available as optional equipment. The engine may also be
equipped with a non-congealing oil cooler.

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.
11-13

11-35.

TROUBLE SHOOTING (Cont).
TROUBLE

HIGH OIL PRESSURE.

LOW OIL TEMPERATURE.

HIGH OIL TEMPERATURE.

11-14

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.
Replace bulb.

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.
passages.

Oil cooler oil passages clogged.

Attempt to drain cooler. 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.

Clean air

11-35.

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.

11-36.

FULL-FLOW OIL FILTER.

11-37. DESCRIPTION. An external oil filter may
be installed on the engine. The filter and filter adapter replace the regular engine oil pressure screen.
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 filter element
should become clogged.
11-38. REMOVAL AND INSTALLATION.
figure 11-4.)

(Refer to

NOTE
Filter element replacement kits are available from the Cessna Service Parts Center.
a. Remove engine cowling in accordance with paragraph 11-3.
b. Remove both safety wires from filter can and
unscrew hollow stud (1) to detach filter assembly
from adapter (10) 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 filter can (4) and discard lower
gasket (6).
e. Pull filter element (5) out of filter can (4).
NOTE
Before discarding removed filter element (5),
remove the outer perforated paper cover;
using a sharp knife, 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 filter can (4)
in solvent and dry with compressed air.
NOTES
When installing a new filter element (5), it
is important that all gaskets are clean, lubricated and positioned properly, and that
the correct amount of torque is applied to
the hollow stud (1). If the stud is undertorqued, oil leakage will occur. If the stud
is over-torqued, the filter can might possibly be deformed, again causing oil leakage.
· Lubricate all rubber grommets in the new
filter element, lid gaskets and metal gasket with clean engine oil or general purpose
grease before installation. Dry gaskets
may cause false torque readings, again
resulting in oil leakage.

11-15

NOTE
Do NOT substitute automotive gaskets for any gaskets
used in this assembly. Use only approved gaskets
listed in the Parts Catalogs.

NUT DELETED ON CURRENT
INSTALLATIONS (DISCARD AT
NEXT FILTER ELEMENT
CHANGE)

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

Hollow Stud
Metal Gasket
Safety Wire Tabs
Can
Filter Element
Lower Gasket
Lid
Nut
Upper Gasket

10.

Adapter

11.
12.
13.
14.
15.

Adapter Nut
O-Ring
Bypass Valve
Plug
Thread Insert

4
3

Figure 11-4.
11-16

Full Flow Oil Filter

1/2
1/2 " (TYP)

1/2

240

(TYP)
3/4"
1-11/16 "R --

1-7/32 "

Figure 11-5.

Oil Filter Adapter Wrench Fabrication

* Before assembly, place a straightedge across
bottom of filter can. Check for distortion or
out-of-flat condition greater than 0.010 inch.
Install a new filter can if either of these conditions exist.
* After installing a new gasket on lid, turn lid
over. If gaskets falls, try a different gasket
and repeat test. If this gasket falls off, install a new lid.
g. Inspect the adapter gasket seat for gouges, deep
scratches, wrench marks and mutilation. If any of
these conditions are found, install a new adapter.
h. Place a new filter element (5) in can (4) and insert the hollow stud (1) with a new metal gasket (2)
in place, through the filter can and element.
i. Position a new gasket (6) inside flange of lid (7)
and place lid in position on filter can.
j. With new gasket (9) on face of lid, install filter
can assembly on adapter (10). While holding filter
can to prevent turning, tighten hollow stud (1) and
torque to 20-25 lb-ft (240-300 lb-in), using a torque
wrench.
k. Install all parts removed for access and service
the engine with the proper grade and quantity of engine oil. One additional quart of oil is required each
time the filter element is changed.
1. Start engine and check for proper oil pressure.
Check for oil leakage after warming up the engine.
m. Again check for oil leakage after engine has been
run at high power setting (preferably a flight around
the field).

n. Check to make sure filter can has not been making contact with any adjacent parts due to engine
torque.
o. While engine is still warm, recheck torque on
hollow stud (1) then safety stud to lower tab (3) on filter can and safety adapter (10) to upper tab on filter
can.
11-39.

FILTER ADAPTER.

11-40. REMOVAL. (Refer to figure 11-4.)
a. Remove filter assembly in accordance with paragraph 11-38.
NOTE
A special wrench adapter for adapter nut
(11) (Part No. SE-709) is available from
the Cessna Service Parts Center, or one
may be fabricated as shown in figure 11-5.
Remove any engine accessory that interferes with removal of the adapter.
b. Note angular position of adapter (10), then remove safety wire and loosen adapter nut (11).
c. Unscrew adapter and remove from engine. Discard adapter O-ring (12).
11-41. DISASSEMBLY, INSPECTION AND REASSEMBLY. Figure 11-4 shows the relative position of the
internal parts of the filter adapter and may be used
as a guide during installation of parts. The bypass
valve is to be installed as a complete unit, with the
11-17

valve being staked three places. The heli-coil type
insert (15) 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 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.
11-42. INSTALLATION.
a. Assemble adapter nut (11)and new O-ring (12)
on adapter (10) in sequence illustrated in figure 11-4.
b. Lubricate 0-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 engine until O-ring seats against
engine boss without turning adapter nut (11). 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. Maintaining this position of the adapter,
tighten adapter nut to 50-60 lb-ft (600-700 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 11-38. Be sure to service the
engine oil system.
11-43.

OIL COOLER.

11-44. DESCRIPTION. A non-congealing oil cooler
may be installed on the aircraft. The cooler is
mounted on the right forward side of the engine crankcase directly in front of number five cylinder and has
no external oil lines. 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
oil coolers, 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.
11-45.

ENGINE FUEL SYSTEM.

11-46. DESCRIPTION. The engine is equipped with
a carburetor mounted at the lower side of the engine.
The carburetor is of the plain-tube fixed-jet type and
has such features as an enclosed accelerating pump
mechanism, simplified fuel passages to prevent vapor
locking, idle cut-off to prevent starting of the engine
accidentally and manual mixture control for leaning.
For overhaul and repair of the carburetor, refer to
the manufacturer's overhaul and repair manual.

11-18

Change 1

11-47.

CARBURETOR.

11-48. REMOVAL AND INSTALLATION.
a. Place fuel selector valve in the OFF position.
b. Remove engine cowling in accordance with paragraph 11-3.
c. Drain fuel from strainer and lines with strainer
drain control.
d. Remove airbox in accordance with paragraph
11-53.
e. Disconnect throttle and mixture controls at
carburetor. Note EXACT position, size and number
of attaching washers and spacers for reference on
reinstallation.
f. Disconnect and cap or plug fuel line at carburetor.
g. Remove safety wire, nuts and washers attaching
carburetor to intake manifold and remove carburetor
and mounting gasket.
h. Reverse the preceding steps for reinstallation.
Use new gaskets when installing carburetor. Rig
controls in accordance with paragraphs 11-73, 11-74
and 11-75.(Check carburetor throttle arm to idle stop
arm attachment for security and proper safetying at
each normal engine inspection inaccordance with figure 11-8.)
11-49. IDLE SPEED AND MIXTURE ADJUSTMENTS.
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 (idle rpm) 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.
b. Advance throttle to increase engine speed to
approximately 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 screw at upper end of
carburetor intake throat to obtain a slight and momentary gain of 25 rpm maximum at 1000 rpm engine speed as mixture control is moved from full IN
(RICH) otward idle cut-off position. Return control
to full IN (RICH) to prevent engine stoppage.
e. If mixture is set too LEAN, engine speed will
stop immediately, thus requiring a richer mixture.

Turn adjusting screw OUT (counterclockwise) for a
richer mixture.
f. If mixture is set too RICH, engine speed will
increase above 25 rpm, thus requiring a leaner
mixture. Turn adjusting screw IN (clockwise) 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.
11-50. INDUCTION AIR SYSTEM.
11-51. DESCRIPTION. Ram air enters the induction air system through a filter at the front of the
lower cowling and is ducted to the airbox at the carburetor. From the induction airbox the filtered air
is directed to the inlet of the carburetor, mounted
on the lower side of the engine, through the carburetor, where fuel is mixed with the air, to the intake
manifold. From the intake manifold, the fuel-air
mixture is distributed to each cylinder by separate
intake pipes. The intake pipes are attached to the
manifold with hoses and clamps and to the cylinder
with a four bolt flange sealed with a gasket. A butterfly valve, located in the airbox, may be operated
manually from the cabin to permit the selection of
either cold or heated air. When the induction air door
is closed, heated air is drawn from a shroud on the
left exhaust stack assembly.
11-52. AIRBOX.
11-53. REMOVAL AND INSTALLATION.
a. Remove engine cowling in accordance with paragraph 11-3.
b. Disconnect flexible duct from left side of airbox.
c. Disconnect boot from forward end of airbox.
d. Disconnect carburetor heat control at arm on
right side of airbox and remove clamp securing control to airbox.

e. Remove mounting bolt safety wire, remove bolts
and gasket and carefully remove airbox.
f. Reverse the preceding steps for reinstallation.
Rig carburetor heat control in accordance with paragraph 11-75.
11-54. CLEANING AND INSPECTION. Clean metal
parts of the induction air box 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 air
box. Inspect gaskets and install new gaskets, if damaged. Check manually-operated air door for ease of
operation and proper rigging.
11-55.

INDUCTION AIR FILTER.

11-56. DESCRIPTION. An induction air filter,
mounted at the induction air inlet on the front of the
lower cowling, removes dust particles from the ram
air entering the engine.
11-57. REMOVAL AND INSTALLATION.
a. (THRU AIRCRAFT SERIALS 18261425 AND ALL
A182 AIRCRAFT.) Release the quick-release fasteners securing filter assembly and lift filter out of nose
cap.
b. (BEGINNING WITH 18261426.) Remove screws
securing filter cover, release the quick-release
fasteners securing filter assembly and lift filter out
of nose cap.
c. Reverse the preceding steps for reinstallation.
11-58. CLEANING AND INSPECTION. Clean and
inspect filter in accordance with instructions in
Section 2.
NOTE
If air filter gasket becomes loose,
bond with EC-1300L or equivalent.
11-59.

IGNITION SYSTEM.

11-60. DESCRIPTION. The ignition system is comprised of two magnetos, two spark plugs in each cylinder, an ignition wiring harness, an ignition switch

Change 3

11-19

Figure_, 11C-6.

Schem

LEFT

RIGHT

FIRING ORDER 1-6-3-2-5-4

Figure 11-6.
11-20

SPARK PLUGS

Ignition Schematic

11-61.

TROUBLE SHOOTING.
TROUBLE

ENGINE FAILS TO START.

ENGINE WILL NOT
IDLE OR RUN PROPERLY.

ENGINE WILL NOT IDLE
OR RUN PROPERLY (Cont).

PROBABLE CAUSE

REMEDY

Defective ignition switch.

Check switch continuity.
if defective.

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 11-67.

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 11-67.

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.

Replace

11-21

11-62.

MAGNETOS.

11-63. DESCRIPTION. 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
have timing marks, visible through the air vent holes,
for timing to the engine. A timing hole is provided
in the bottom of the magneto adjacent to the magneto
flange. 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 or internal
timing may be disturbed.
11-64. REMOVAL.
a. Remove engine cowling in accordance with paragraph 11-3.
b. Tag for identification and remove high tension
wires from the magneto being removed.

IWARNING
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,

11-22

Change 3

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 dislodged
from the gear hub and fall into the engine.
NOTE
For inspection of impluse coupling on aircraft serials 18256685 THRU 18263179
refer to Cessna Single-engine Service
Letter SE74-21, dated September 27, 1974.
11-65. 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 in the bottom of the magneto next to the flange and 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
plug holes are approximately aligned.
NOTE
The side of the magneto with the manufacturer's insignia has a red timing mark
and the side opposite to the insignia has
a black timing mark viewed through the
vent plug holes. The distributor gear
also has a red timing mark and a black
timing mar. 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 aligned for 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.
11-66. 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 11-12 for the advanced firing position of number one piston.

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 magneto 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.

the magneto drive gear out of mesh with its
drive gear and rotate it to the aligned angle,
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

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) and insert a timing
pin (or 0. 093 inch 6-penny nail) through the timing
hole in the bottom of the magneto next to the flange
and into the mating hole in the rotor shaft. This
locks the magneto approximately in the firing position while installing on the engine.
NOTE
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

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
11-12.
k. Connect toggle switch (primary) lead to the capacitor terminal on the magneto.
1. Inspect magneto installation and install engine
cowling in accordance with paragraph 11-3.
11-67. 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
inspection. If timing is 22 ° (plus zero, minus 2°),
Change 1

11-23

THESE CONTACT POINTS ARE USABLE

Figure 11-7.

THESE CONTACT POINTS NEED REPLACEMENT

Magneto Contact Breaker Points

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 magneto-to-engine
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 within
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 11-7 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.
NOTE
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.

11-24

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
brush for moisture.
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 surfaces 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.
4. Check the carbon brush on the distributor
gear for excessive wear. The brush must extend a
minimum of 1/32 inch beyond the end of the gear
shaft. The spring which the carbon brush contacts
should be bent out approximately 20 degrees from
vertical, since spring pressure on the brush holds
the distributor gear shaft against the thrust bearing
in the distributor cap.
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.

to reduce RPM drop on single ignition. 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 has
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"
position.
f. The rpm drop should not exceed 150 rpm on
either magneto or show greater than 50 rpm differential between magnetos. A smooth rpm drop-off
past normal is usually a 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

lower spark plugs is usually more rapid than
that of the upper spark plugs, rotating helps
prolong spark plug life.
11-70.

ENGINE CONTROLS.

11-71. DESCRIPTION. The throttle, mixture, propeller and carburetor heat controls are of the pushpull 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 for more or less friction as desired. The
friction knob prevents vibration induced "creeping" of
the control. The carburetor heat control has no locking device.
NOTE
Some controls have intricate parts that will
fall out and possibly be lost if the control is
pulled from the housing while it is disconnected.
11-72. 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.

will usually confirm whether a deficiency exists.
NOTE
An absence of rpm drop may be an indication
of faulty grounding of one side of the ignition
system, a disconnected ground lead at magneto or possibly the magneto timing is set
too far in advance.
11-69. SPARK PLUGS. Two spark plugs are installed in each cylinder and screw into helicoil 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
At each 100-hour inspection, remove, clean,
inspect and regap all spark plugs. Install

Some engine controls have a small retaining
ring brazed (or attached with epoxy resin)
near the threaded end (engine end) of the control. The purpose of these retaining rings is
to prevent inadvertent withdrawal of and possible damage to the knob end of the controls
while jam nuts and rod ends are removed.
* 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.
11-73.

THROTTLE CONTROL.
NOTE

Before rigging throttle control shown in figure 11-8, check that staked connection (4)
between rigid conduit (2) and flexible conduit
(3) is secure. If any indication of looseness
or breakage is apparent, replace the throttle
control before continuing with the rigging
procedure.
11-25

10-32 Bolt and Lock
Nut, Torque to 35-

3.
4.
5.
6.
7.

Figure 11-8.

NOTE
Refer to the inspection chart in Section 2
for inspection and/or replacement interval
for the throttle control.
11-74. MIXTURE CONTROL.
a. Push mixture control full in, then pull it out approximately 1/8 inch for cushion.
b. Loosen clamp securing the control to the engine.
c. Shift control housing in the clamp so that the
mixture arm on the carburetor is in the full open position (RICH). Tighten the clamp in this position.
d. Unlock and pull mixture control full out. Check
that idle mixture arm on carburetor is full closed
(IDLE CUT-OFF).
e. Check that the bolt and nut at the mixture arm
on carburetor secures the control wire and that the
Change 1

Safety Wire

Throttle Control and Throttle Arm to Idle Stop Attachment

a. Pull throttle control out (idle position) and remove throttle control knob (1).
b. Screw jam nut (7) all the way down (clockwise)
and install throttle knob. Screw the knob securely
against the jam nut. Do not back jam nut out. This
will prevent bottoming and possible damage to the
staked connection.
c. Disconnect throttle control at the carburetor
throttle arm, push throttle control in until jam nut
hits friction lock (6) while the friction lock is loose,
then pull control out approximately 1/8 inch for cushion. Note position of large washer at carburetor end
of control. Install washer in same position when connecting control to arm.
d. Tighten friction lock (6), being careful not to
change position of the throttle.
e. Move throttle arm on carburetor to full open,
adjust rod end at end of throttle control to fit and
connect to arm on carburetor.
f. Release friction lock and check full travel of
arm on carburetor. If further adjustment is required, make all adjustment at the carburetor end
of control. DO NOT change jam nut (7) setting.
g. Tighten rod end locknuts at carburetor end of
control. Be sure to maintain sufficient thread engagement between rod end and control.

11-26

Flexible Conduit
Staked Connection
Instrument Panel
Friction Lock
Jamb Nut

bolt will swivel in the arm.
f. Bend the wire tip 90 degrees to prevent it from
being withdrawn if the attaching nut should become
loose.
g. When installing a new control, it may be necessary to shorten the wire and/or control housing.
h. The mixture arm on the carburetor must contact
the stops in each direction, and the control should
have approximately 1/8 inch cushion when pushed in.
NOTE
Refer to the inspection chart in Section 2
for inspection and/or replacement interval
for the mixture control.

11-75. CARBURETOR HEAT CONTROL.
a. Loosen clamp securing the control to the bracket
on the airbox.
b. Push control full in, then pull it out approximately 1/8 inch from panel for cushion.
c. Shift control housing in its clamp so that the
valve in the airbox is seated in the full open position.
Tighten clamp in this position.
d. Pull out on the control and check that the air
valve inside the airbox seats in the opposite direction.
e. Check that bolt and nut on the air valve lever
secures the control wire and that the bolt will swivel
in the lever.
f. Bend the wire tip 90 degrees to prevent it from
being withdrawn if the attaching nut should become
loose.
NOTE
Refer to the inspection chart in Section 2
for inspection and/or replacement interval
for the carburetor heat control.
11-76.
13.

PROPELLER CONTROL.

11-77.

STARTING SYSTEM.

Refer to Section

11-78. 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 overrunning clutch in the starter adapter, which incorporates worm reduction gears. The starter motor is
located just aft of the right rear cylinder.

11-79.

CAUTION
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

Replace

Change 1

11-27

11-80. PRIMARY MAINTENANCE. The starting
circuit should be inspected at regular intervals, the
frequency of which should be determined by the
amount of service and conditions under which the
equipment is operated. Inspect the battery and wiring. Check battery for fully charged condition, proper electrolyte level with approved water and terminals for cleanliness. Inspect wiring to be sure that
all connections are clean and tight and that the wiring
insulation is sound. Check that the brushes slide
freely in their holders and make full contact on the
commutator. When brushes are worn to one-half of
their original length, install new brushes (compare
brushes with new brushes). Check the commutator
for uneven wear, excessive glazing or evidence of
excessive arcing. If the commutator is only slightly
dirty, glazed or discolored, it may be cleaned with a
strip of No. 00 or No. 000 sandpaper. If the commutator is rough or worn, it should be turned in a lathe
and the mica undercut. Inspect the armature shaft
for rough bearing surfaces. New brushes should be
properly seated when installing by wrapping a strip
of No. 00 sandpaper around the commutator (with
sanding side out) 1-1/4 to 1-1/2 times maximum.
Drop brushes on sandpaper covered commutator and
turn armature slowly in the direction of normal rotation. Clean sanding dust from motor after sanding
operations.
11-81. STARTER MOTOR.
11-82. REMOVAL AND INSTALLATION.
a. Remove engine cowling in accordance with paragraph 11-3.

When disconnecting starter electrical cable,
do not permit terminal bolt to rotate. Rotation of the bolt could break the conductor
between bolt and field coils causing the
starter to be inoperative.
b. Disconnect battery cables and insulate as a
safety precaution.
c. Disconnect electrical cable at starter motor.
d. Remove nuts and washers securing motor to
starter adapter and remove motor. Refer to engine
manufacturer's overhaul manual for adapter removal.
e. Reverse the preceding steps for reinstallation.
Install a new O-ring seal on motor, then install motor.
Be sure motor drive engages with the adapter drive
when installing.
11-83.

EXHAUST SYSTEM.

11-84. 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. Each exhaust
stack assembly connects to the muffler beneath the
engine. The muffler is enclosed in a shroud which
captures exhaust heat which is used to heat the cabin.
11-28

Change 3

The tailpipe is welded to the muffler. A shroud is
attached to the left exhaust stack to provide heated
air for the carburetor heat source.
11-85. REMOVAL AND INSTALLATION. (Refer to
figure 11-9.)
a. Remove engine cowling in accordance with paragraph 11-3.
b. Disconnect ducts from heater shroud on muffler
assembly.
c. Disconnect duct from shroud on left exhaust
stack assembly.
d. Remove nuts, bolts and clamps attaching stack
assemblies to the muffler.
e. Loosen nuts attaching exhaust stacks to the
cylinders and remove muffler assembly.
f. Remove nuts attaching exhaust stack assemblies
to the cylinders and remove exhaust stacks and gaskets.
g. Reverse the preceding steps for reinstallation.
Install a new copper-asbestos gasket between each
riser and its mounting pad on each cylinder, regardless of apparent condition of those removed. Torque
exhaust stack nuts at cylinders to 100-110 poundinches.
11-86. INEPECTION. Since exhaust systems of this
type are subject to burning, cracking and general deterioration from alternate thermal stresses and vibrations, inspection is important and should be accomplished every 100 hours of operation. Also, a thorough inspection of the engine exhaust system should
be made to detect cracks causing leaks which could
result in loss of engine power. To inspect the engine
exhaust system, proceed as follows:

inspected.
inspected.
NOTE
Especially check the areas adjacent to welds
and slip joints. Look for gas deposits in surrounding areas, indicating that exhaust gases
are escaping through a crack or hole or around
the slip joints.
b. After visual inspection, an air leak check should
be made on the exhaust system as follows:
1. Attach the pressure side of an industrial
vacuum cleaner to the tailpipe opening, using a rubber plug to effect a seal as required.
NOTE
The inside of vacuum cleaner hose should be
free of any contamination that might be blown
into the engine exhaust system.
2. With vacuum cleaner operating, all joints
in the exhaust system may be checked manually by
feel, or by using a soap and water solution and
watching for bubbles. Forming of bubbles is considered acceptable, if bubbles are blown away
system is not considered acceptable.
c. Where a surface is not accessible for a visual
inspection, or for a more positive test, the following
procedure is recommended.

BEGINNING WITH
18264231

.38 " minimum clearance between exhaust
muffler and induction air duct assembly.

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

Clamp Half
Exhaust Stack Assembly
Riser
Cabin Heat Outlet
Cabin Heat Inlet
Shroud
Muffler
Tailpipe

Detail
Detail B
BEGINNING WITH 18264231

Figure 11-9.

Exhaust System
Change 3

11-29

1. Remove exhaust stack assemblies.
2. Use rubber expansion plugs to seal openings.
3. Using a manometer or gage, apply approximately 1-1/2 psi (3 inches of mercury) air pressure
while each stack assembly is submerged in water.
Any leaks will appear as bubbles and can be readily
detected.
4. It is recommended that exhaust stacks found
defective be replaced before the next flight.
d. After installation of exhaust system components
perform the inspection in step "b" of this paragraph
to ascertain that system is acceptable.

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
result 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.

11-87.

EXTREME WEATHER MAINTENANCE.

11-88. COLD WEATHER. Cold weather starting
will be made easier by the installation of an oil dilution system, an engine primer system and a ground
service receptacle. The primer system is manuallyoperated 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 11-92 for use of
the external power receptacle.
The following may also be used to assist engine starting in extreme cold weather. After the last flight of
the day, drain the engine oil into a clean container so
the oil can be preheated. Cover the engine to prevent
ice or snow from collecting inside the cowling. When
preparing the aircraft for flight or engine runup after
these conditions have been followed, preheat the drained engine oil.

IWARNING
Do not heat the oil above 121°C (250°F). A
flash fire may result. Before pulling the
propeller through, ascertain that the magneto switch is in the OFF position to prevent
accidental firing of the engine.
After preheating the engine oil, gasoline may be
mixed with the heated oil in a ratio of 1 part gasoline
to 12 parts engine oil before pouring into the engine
oil sump. If the free air temperature is below minus
29°C (-20°F), the engine compartment should be preheated by a ground heater. After the engine compartment has been preheated, inspect all engine drain and
vent lines for presence of ice. After this procedure
has been complied with, pull propeller through several revolutions by hand before attempting to start the
engine.
CAUTION
Due to the desludging effect of the diluted
oil, engine operation should be observed
closely during the initial warm-up of the
engine. Engines that have considerable
amount of operational hours accumulated
since their last dilution period may be
seriously affected by the dilution process.
This will be caused by the diluted oil dislodging sludge and carbon deposits within
11-30

Change 3

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.
11-89. HOT WEATHER. Engine mis-starts characterized by weak, intermittent explosions followed by
puffs of black smoke from the exhaust are caused by
over-priming or flooding. This situation is more apt
to develop in hot weather or when the engine is hot.
If it occurs, repeat the starting routine with the throttie approximately one-half OPEN and the mixture control in IDLE CUT-OFF. As the engine fires, move
the mixture control to full RICH and decrease the
throttle to desired idling speed.
Engine mis-starts characterized by sufficient power
to disengage the starter but dying after 3 to 5 revolutions are the result of an excessively lean mixture
after the start. This can occur in either warm or
cold temperatures. Repeat the starting routine with
additional priming.

CAUTION
Never operate the starting motor more than
12 seconds at a time. Allow starter motor
to cool between cranking periods to avoid
overheating. Longer cranking periods will
shorten the life of the starter motor.
11-90. 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.
11-91. DUSTY AREAS. Dust induced into the intake
system of the engine is probably the greatest single
cause of early engine wear. When operating in high
dust conditions, service the induction air filter daily
as outlined in Section 2. Also change engine oil and
lubricate airframe items more often than specified.
11-92. GROUND SERVICE RECEPTACLE. With
the ground service receptacle installed, the use of
an external power source is recommended for cold
weather starting, low battery starting and lengthy
maintenance of the aircraft electrical system. Refer
to Section 16 for additional information.
11-93. HAND-CRANKING. A normal hand-cranking
procedure may be used to start the engine.

SECTION 12

FUEL SYSTEM

Page

TABLE OF CONTENTS
.
FUEL SYSTEM .............
Description ..............
Precautions .............
.
......
Trouble Shooting ....
. . . . ...
Fuel Cells . . . . . ..
.
. .....
Description ..
General Precautions .......
Removal .............
.
Repair ............
.
.......
Installation ..
.
Fuel Quantity Transmitters ...
Description ............
Removal and Installation ......

12-1
12-1
12-1
12-2
12-5
12-5
12-5
12-5
12-5
12-10
12-10
12-10

FUEL SYSTEM.

12-1.

12-2. DESCRIPTION. A rubberized bladder-type
fuel cell is located in the inboard bay of each wing.
Fuel is gravity-fed from the cells through the finger
strainers, selector valve and fuel strainer to the
carburetor. Positive ventilation is provided by a
vent line and check valve assembly located in the left
wing cell. The vent line from the check valve assembly extends overboard through the lower wing skin
adjacent to the left wing strut. The fuel supply line
from the lower forward corner of each cell serves
as a combination fuel feed and vapor return line and
is teed into the cell crossover vent line. The strainer is equipped with a quick-drain valve which provides a means of draining trapped water and sediment
from the fuel system.
12-3.

PRECAUTIONS.
NOTE

There are certain general precautions and
rules concerning the fuel system which
should be observed when performing the operations and procedures in this section.
These are as follows:

..........
Fuel Vents ...
Description ............
Checking .............
........
Fuel Selector Valve .
. . . . .
Description ...
Removal and Installation ..
Fuel Strainer ......
...
Description
Removal and Installation ..
Disassembly and Assembly .
Priming System . ........
Description .........
Removal and Installation ...

. 12-10
12-10
12-10
. 12-10
. . . . 12-10
. 12-10
..
.12-11
12-11
. 12-11
. .
12-11
....
. 12-14
. 12-14
. 12-14

a. During all fueling, defueling, purging, repairing
or disassembly, ground the aircraft to a suitable
ground stake.
b. Residual fuel draining from lines and hose constitutes a fire hazard. Use caution to prevent the
accumulation of fuel when lines or hose are disconnected.
c. Cap open lines and cover connections to prevent
thread damage and the entrance of foreign matter.
NOTE
Throughout the aircraft fuel system, from
the fuel cells to the carburetor, use NS-40
RAS-4 (Snap-On Tool Corp., Kenosha, Wisconsin). MIL-T-5544 (Thread Compound),
Antiseize, Graphite-Petrolatum) or equivalent, as a thread lubricant or to seal a leaking connection. Apply sparingly to male
fittings only. omitting the first two threads.
Always ensure that a compound. the residue
from a previously used compound, or any
other foreign material cannot endter the
system.

Change 1

12-1

12-4.

TROUBLE SHOOTING.
TROUBLE

NO FUEL FLOW TO
CARBURETOR

FUEL STARVATION AFTER
STARTING

NO FUEL QUANTITY
INDICATION

SHOP NOTES:

12-2

PROBABLE CAUSE

REMEDY

Fuel selector valve not turned on.

Turn valve on.

Fuel cells empty.

Service with proper grade and
amount of fuel.

Fuel line disconnected or broken.

Connect or repair fuel lines.

Fuel cell outlet screens plugged.

Remove and clean screens and
flush out fuel cells.

Defective fuel selector valve.

Repair or replace selector valve.

Inlet elbow or inlet screen
in carburetor plugged.

Clean or replace.

Plugged fuel strainer.

Remove and clean strainer and screen.

Fuel line plugged.

Clean or replace fuel line.

Partial fuel flow from the preceding causes.

Use the preceding remedies.

Plugged fuel vent.

Refer to paragraph 12-22.

Water in fuel.

Drain fuel cell sumps, lines
and strainer.

Fuel cell 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 15.

Replace if defective.

FUEL QUANTITY INDICATORS

FILLER

...
FUEL CELL SUMP
DRAIN PLUG/VALVE--

III IFUEL

.........

Figure 12-1.

...

CELL SUMP
..........
DRAIN PLUG/VALVE

Fuel System Schematic
12-3
12-3

REFER TO FIGURE 12-3

NOTE

REFER TO FIGURE 12-67

Figure 12-2.

12-4

LONG-RANGE INSTALLATIONS ONLY

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

Hose
Fuel Strainer
Line
Primer Line
Primer
Hose

6.
7.
7.
8.
8.
9.
10.
11.

Finger Strainer
Fuel Filler Cap
Fuel Quantity Transmitter
Crossover Vent Line
Fuel Vent Valve
Vent Line

Figure 12-2. Fuel System
12-4

All fuel hoses should be replaced at engine overhaul or
after
whichever
years, whichever
after 55 years,
first.
comes
comes first.

Fuel System

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

Placard
Valve
Fuel Selector Valve
Gear and Shaft Assembly
Strainer Drain Control
Line
Drain Line
Drain

12-5.

FUEL CELLS.

(RUBBERIZED)

12-9.

12-6. DESCRIPTION. Rubberized, bladder-type
fuel cells are installed in the inboard bay of each
wing panel. These cells are secured by fasteners to
prevent collapse of the flexible cells.
12-7. GENERAL PRECAUTIONS. When storing,
inspecting or handling rubberized, bladder-type fuel
cells, the following precautions should be adhered to:
a. Fold cells as smoothly and lightly as possible
with a minimum number of folds. Place protective
wadding between folds.
b. Wrap cell in moisture-proof paper and place in
a suitable container. Do not crowd cell in container.
Use wadding to prevent movement.
c. Stack boxed cells to allow access to oldest cell
first. Do not allow stacks to crush bottom boxes.
Leave cells in boxes until used.
d. Storage area must be cool, +30°F to 85 ° , and
free of exposure to sunlight, dirt and damage.
Used cells must be cleaned with soap and warm
water prior to storage. Dry and package as outlined
in the preceding steps.
f. Do not carry cells by fittings. Maintain original
cell contours or folds when refolding for boxing.
12-8. FUEL CELL REMOVAL.
a. Drain fuel from applicable cell.
NOTE
Prior to removal of cell, drain fuel, purge
with fresh air, and swab out to remove all
traces of fuel.
b. Remove wing root fairings and disconnect fuel
lines at wing root.
c. Remove clamps from forward and aft fuel cell
bosses at wing root and carefully work fuel strainers
and lines from cell bosses.
d. Disconnect electrical lead and ground strap from
fuel quantity transmitter and carefully work transmitter from fuel cell and wing rib.
e. Remove screws attaching drain adapter to lower
surface of wing.
f. Remove clamps attaching crossover vent line to
fuel cells and work vent line out of cell being removed.
In aircraft equipped with long-range cells, remove
vent extension tube from inside cell. Vent extension
tube is attached to the crossover vent bars on the
cell.
g. Remove fuel filler adapter and gaskets by removing screws attaching adapter to wing and fuel
cell. On aircraft equipped with long-range cells,
remove cover plate and gaskets, and remove nylon
vent tube from inside cell.
h. Working through filler neck opening, loosen
snap fasteners. Tilt snap fasteners slightly when
pulling cell free, to prevent tearing rubber.
i. Collapse and carefully fold cell for removal,
then work cell out of fuel bay through filler opening
in upper wing surface. Use care when removing to
prevent damage to cell.
j. Unfold cell and remove fittings, snap fasteners
and fuel sump drain adapter.

FUEL CELL REPAIR.
NOTE

For fuel cell repair information, refer
to Cessna Service News Letter dated
August 28, 1970. For minor repair, a
fuel cell repair kit is available from
Goodyear, complete with required
materials and instructions.
12-10.

Deleted.

12-11.

Deleted.

12-12.

Deleted.

12-13.

Deleted.

12-14.

Deleted.

12-15.

Deleted.

12-16. FUEL CELL INSTALLATION.
a. Cell compartment must be thoroughly cleaned of
all filings, trimmings, loose washers, bolts, nuts,
etc.
b. All sharp edges of cell compartment must be
rounded off and protective tape applied over any
other sharp edges and protruding rivets.
c. Inspect cell compartment just prior to installation of a cell for conditions noted in the preceding
steps.
d. Install fuel drain adapter and snap fasteners.
e. Check to ensure cell is warm enough to be flexible and fold as necessary to fit through fuel cell access opening.
f. Place cell in compartment, develop it out to full
size and attach fasteners, then reverse procedures
outlined in preceding paragraph for installation. Install all new gaskets when installing cell.
g. On aircraft equipped with long-range cells, install nylon vent tube inside cell, inserting tube
through four hangers in top of cell. If a replacement
cell is being installed, use nylon vent tube removed
from old cell and/or order tube from applicable
Parts Catalog.
h. When tightening screw-type clamps, apply a
maximum of 20 pound-inches torque to clamp screws.
No oil is to be applied to fittings prior to installation.
i. When installing filler adapter, cover plate and
fuel quantity transmitter to the wing and fuel cell,
tighten attaching screw evenly. The sealing or compression surfaces must be assembled when absolutely dry (NO SEALING PASTE IS TO BE USED).
j. After installation has been completed, cell should
be inspected for final fit within compartment, making
certain that cell is extended out to the structure and
no corners are folded in.
k. The final inspection, prior to closing the cell,
should be a close check to ensure that cell is free of
foreign matter such as lint, dust, oil or any installation equipment. If a cell is not thoroughly clean, it
should be cleaned with a lint-free cloth, soaked in
water, alcohol or kerosene. NO OTHER SOLVENT
SHALL BE USED.

(Pages 12-6 and 12-7 Deleted)
Change 1

12-5

Hinge for vent valve (11) must be at top. Tube for vent extends
into fuel cell, then is offset upward. Vent valve (11) is used in
the left wing fuel cell only.

Detail
-

Detail

Fuel Sampler Cup
(Refer to paragraph 2-20)

DetailD
4
16
1. Plug/Valve
2. Gasket
3. Adapter
4. Clamp
5. Fitting
6. Wing Skin

A
7.
8.
9.
10.
11.

Filler Cap
Vent Line
Grommet
Hose
Vent Valve

12.
13.
14.
15.
16.

Ground Strap
Fuel Quantity
Transmitter
Hanger (Typ)
Strainer
Protector

12
Detail C
FUEL QUANTITY TRANSMITTER
INSTALLATION AND GROUNDING

Figure 12-3.
12-8

Change 2

Fuel Cell Installation (Sheet 1 of 2)

Hinge for vent valve (12) must be at top. Tube for valve extends
into fuel cell, then is offset upward. Vent valve (12) is used in
the left wing fuel cell only.

-4

.DetailB
Detail A

uel Sample Cup
(Refer to paragraph 2-20)

1. Plug/Valve
2. Gasket
3. Adapter
4. Clamp
5. Fitting
6. Wing Skin
7. Cover Plate

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

Filler Cap
Vent Line
Grommet
Hose
Vent Valve
Ground Strap

Figure 12-3.

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

Fuel Quantity
Transmitter
Nylon Tube
Strainer
Protecter
Vent Adapter
Hanger (Typ)

Detail C
FUEL QUANTITY TRANSMITTER
INSTALLATION AND GROUNDING

Fuel Cell Installation (Sheet 2 of 2)
Change 2

12-9

NOTE
Throughout the aircraft fuel system, from
the fuel cells to the carburetor, use NS-40
RAS-4 (Snap-On Tool Corp., Kenosha, Wisconsin), MIL-T-5544 (Thread Compound),
Antiseize, Graphite-Petrolatum) or equivalent. as a thread lubricant or to seal a leaking connection. Apply sparingly to male
fittings only, omitting the first two threads.
Always ensure that a compound, the residue
from a previously used compound, or any
other foreign material cannot enter the
system.
12-17.

FUEL QUANTITY TRANSMITTERS.

b. Blow into tube to slightly pressurize cell. If
air can be blown into cell, vent line is open.
c. After cell is slightly pressurized, insert end of
rubber tube into a container of water and watch for
a continuous stream of bubbles, which indicates the
bleed hole in valve assembly is open and relieving
pressure.
d. After completion of step "c", blow into tube
again to slightly pressurize the cell. Crimp rubber
tube to retain pressure within the cell. Loosen, but
do not remove filler cap on opposite wing to check
cell crossover line. If pressure escapes from
filler cap, crossover line is open. Remove rubber
tube from end of vent line beneath the wing after

completion of check.
NOTE

12-18. DESCRIPTION. Refer to Section 15 for a
complete description of the transmitters.
12-19. REMOVAL AND INSTALLATION.
Section 15 for procedures.
12-20.

Refer to

FUEL VENTS.

12-21. DESCRIPTION. A vent line is installed in
the outboard end of the left fuel cell and extends
overboard through the lower wing skin. The inboard
end of the vent line extends into the fuel cell, then
forward and slightly upward. A vent valve is installed on the inboard end of the vent line inside the fuel
cell, and a crossover line connects the cells together. On aircraft equipped with long-range cells,
a nylon vent tube is attached to the crossover line
at the inboard end of each cell. This vent tube extends into the fuel cell, and is suspended by four
hangers in the top of the cell.
12-22. CHECKING. Field experience has demonstrated that the fuel vent can become plugged, with
possible fuel starvation of the engine or collapse of
the fuel cells. Also, the bleed hole in the vent valve
assembly could possibly become plugged, allowing pressure from expanding fuel to pressurize the
cells. The following procedure may be used to
check the vent and bleed hole in the valve assembly.
a. Attach a rubber tube to the end of vent line beneath the wing.

12-10

Change 1

Remember that a plugged vent line or bleed
hole can cause either fuel starvation and
collapsing of fuel cells or the pressurization of cells by fuel expansion.
e. Any fuel vent found plugged or restricted must
be corrected prior to returning aircraft to service.
NOTE
The fuel vent line protruding beneath the wing
near the wing strut must be correctly aligned
to avoid possible icing of the vent tube. Dimensions are shown in figure 12-4.
12-23.

FUEL SELECTOR VALVE.

12-24. DESCRIPTION. A four position fuel selector
valve is located between the pilot and copilot positions
on the pedestal. The positions on the valve are labeled "OFF, LEFT, BOTH ON and RIGHT. " Valve repair consists of replacement of O-rings and washers.
Figure 12-5 illustrates the proper relationship of
parts and may be used as a guide during disassembly
and assembly.
12-25. REMOVAL AND INSTALLATION. (See figure 12-5.)
a. Completely drain all fuel from cells, lines,

^-S/
^

^L^-^-^

-----

INBOARD

VIEW
LOOKING
FORWARD

Wing

2.
3.
4.

Fairing
Vent
Strut

NOTE
Dimensions must be within +. 03" tolerance.

Figure 12-4.

strainer and selector valve. (Observe precautions
in paragraph 12-3.)
b. Remove selector valve handle.
c. Remove pedestal cover,
d. Remove carpeting as necessary to gain access
to plates at bottom and aft of pedestal.
e. Disconnect handle drive shaft from valve.
f. Disconnect and cap or plug all fuel lines at
valve.
g. Remove screws attaching valve to structure
and remove valve.
h. Reverse the preceding steps for installation.
Prior to installing access plates, service fuel cells
and check for leaks.
12-26.

FUEL STRAINER.

(See figure 12-6.)

12-27. DESCRIPTION. The fuel strainer is mounted at the firewall in the lower engine compartment
and 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 and is accessible
through the oil dipstick door.
NOTE
The fuel strainer can be disassembled,
cleaned and reassembled without removing the assembly from the aircraft.
(Refer to paragraph 12-29.)
12-28. REMOVAL AND INSTALLATION.
ure 12-6.)

(See fig-

Fuel Vent Location

a. Remove cowling as necessary to gain access to
strainer.
b. With selector valve in "OFF" position, drain
fuel from strainer and lines with strainer quickdrain control.
c. Disconnect and cap or plug all fuel lines and
controls from strainer. (Observe precautions in
paragraph 12-3.)
d. Remove bolts attaching assembly to firewall and
remove strainer.
e. Reverse the preceding steps for installation.
With selector valve in "ON" position check for leaks
and proper operation of quick-drain valve.
12-29. DISASSEMBLY AND ASSEMBLY. (See figure 12-6.)
a. With selector valve in "OFF" position, drain
fuel from bowl and lines with quick-drain control.
b. Remove drain tube, safety wire, nut and washer at bottom of filter bowl and remove bowl.
c. Carefully unscrew standpipe and remove.
d. Remove filter screen and gasket. Wash filter
screen and bowl with solvent (Federal Specification
P-S-661, or equivalent) and dry with compressed air.
e. Using a new gasket between filter screen and
top assembly, install screen and standpipe. Tighten
standpipe only finger tight.
f. Using all newO-rings, install bowl. Note that
step-washer at bottom of bowl is installed so that
step seats against O-ring. Connect drain tube.
g. With selector valve in "ON" position, check for
leaks and proper operation of quick-drain valve.
h. Safety wire bottom nut to top assembly. Wire
must have right hand wrap, at least 45 degrees.
12-11

6
NOTE
Do not disassemble selector valve

and alteration of some 1973 and 1974
selector valve gear and shaft assemblies.

Repair kit No. 0716613-200 is available from the Cessna Service Parts
Center for replacement of seals,
washer and O-rings in fuel selector

with Serial 1825399.

2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
18.
19.
20.
21.
22.
23.

Handle
Washer/
Cap
Gear Retainer Assembly
Valve Handle Shaft
Gear
Roll Pin
Valve Shaft
Coupling
Cotter Pin
Fuel Line
Plug
Fuel Selector Valve
Elbow
Body

23
13

22
Detail

0-Ring
Seal
Spring
Lockwasher
Screw
Washer

Figure 12-5.
12-12

Change 2

Fuel Selector Valve

NOTE16

1. Spring
2. Washer
3. Plunger
4. Top
5. Drain Control

6.
7.
8.
9.
10.
11.

Plate
O-Ring
Gasket
Filter
Retainer Ring
Standpipe

Figure 12-6.

12.
13.
14.
15.
16.

0-Ring
Bowl
O-Ring
Nut
Drain Line

Fuel Strainer
12-13

12-30.

PRIMING SYSTEM.

12-31. DESCRIPTION. The priming system is
comprised of a plunger-type manually-operated
primer, which draws fuel from the strainer and
forces it through a tee fitting to the aft end of each
intake manifold. Injecting the fuel into each manifold primes both banks of cylinders,
12-32. REMOVAL AND INSTALLATION.
a. With selector valve in "OFF" position, drain
fuel from strainer and lines with quick-drain control,
b. Disconnect and cap or plug all fuel lines at
primer. (Observe precautions in paragraph 12-3. )
c. Unscrew knurled nut and remove plunger from
pump body.

12-14

Remove pump body from instrument panel.
NOTE
Visually inspect primer lines for crushed,
kinked or broken condition. Ensure proper
clamping to prevent fatigue due to vibration
and chafing.

e. Prior to installing a primer, check for proper
pumping action and positive fuel shut-off in the
locked position.
f. Reverse the preceding steps for installation.
With selector valve in "BOTH" position, check for
leaks and proper pumping action.

SECTION 13
PROPELLERS AND PROPELLER GOVERNORS

Page

TABLE OF CONTENTS
PROPELLERS .............
..............
Description3
Repair ...............
Trouble Shooting ............
|
Removal13 ...........
Installation ..............

13-1.

13-1
13-1
13-1
13-2
13-3
13-3

PROPELLERS.

13-2. DESCRIPTION. The aircraft is equipped with
an all-metal, constant-speed, governor-regulated
propeller. The constant-speed propeller is singleacting, in which engine oil pressure, boosted and
regulated by the governor is used to obtain the correct blade pitch for the engine load. Engine lubricating oil is supplied to the power piston in the propeller hub through the crankshaft. The amount and pressure of the oil supplied is controlled by the enginedriven governor. Increasing engine speed will cause
oil to be admitted to the piston, thereby increasing
the blade pitch. Conversely, decreasing engine speed
will result in oil leaving the piston, thus decreasing
the blade pitch. During the 1969 model year, a new

PROPELLER GOVERNORS .........
Description ..............
Trouble Shooting ............
Removal. ...............
Installation ..............
High-RPM Stop Adjustment .......
Rigging Propeller Governor Control

13-3
13-3
13-5
13-5
13-5
13-5
. . . 13-6

threadless blade propeller is installed. With this
type blades, the propeller balance weights are moved
to a bracket on the propeller cylinder nearer the center line of the propeller. Figure 13-1 illustrates the
different propellers used on the aircraft.
13-3. REPAIR. Metal propeller repair first involves
evaluating the damage and determining whether the
repair will be a major or minor one. Federal Aviation Regulations, Part 43 (FAR 43), and Federal
Aviation Agency, Advisory Circular No. 43. 13 (FAA
AC No. 43.13), define major and minor repairs, alterations and who may accomplish them. When making repairs or alterations to a propeller FAR 43,
FAA AC No. 43. 13 and the propeller manufacturer's
instructions must be observed.
Change 1

13-1

13-4.

TROUBLE SHOOTING.
TROUBLE

FAILURE TO CHANGE PITCH.

PROBABLE CAUSE

REMEDY

Governor control disconnected or
broken.

Check visually. Connect or replace control.

Governor not correct for
propeller. (Sensing wrong.)

Check that correct governor is
installed. Replace governor.

Defective governor.

Refer to paragraph 13-9.

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 13-9.

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 11 for

ENGINE SPEED WILL NOT
STABILIZE.

Sludge in governor.

Refer to paragraph 13-9.

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.

FAILURE TO CHANGE PITCH
FULLY.

SHOP NOTES:

13-2

Change 1

13-4.

TROUBLE SHOOTING (Cont.)
TROUBLE

PROBABLE CAUSE

OIL LEAKAGE AT PROPELLER MOUNTING FLANGE.

OIL LEAKAGE AT ANY
OTHER PLACE.

REMEDY

Damaged O-ring and seal between
engine crankshaft flange and
propeller.

Check visually. Remove propeller
and install O-ring seal.

Foreign material between
engine crankshaft flange and
propeller mating surfaces or
mounting nuts not tight.

Remove propeller and clean
mating surfaces; install new
O-ring and tighten mounting
nuts evenly to torque value
in figure 13-1.

Defective seals, gaskets,
threads, etc., or incorrect
assembly.

Propeller repair or replacement
is required.

13-5. REMOVAL. (Refer to figure 13-1.)
a. Remove spinner attaching screws and remove
spinner (1), spinner support (2) and spacers (3).
Retain spacers (3).
b. Remove cowling as required for access to
mounting nuts (14).
c. Loosen all mounting nuts (14) approximately
1/4 inch and pull propeller (6) forward until stopped
by nuts.
NOTE
As the propeller (6) is separated from the
engine crankshaft flange, oil will drain
from the propeller and engine cavities.
d. Remove all propeller mounting nuts (14) and
pull propeller forward to remove from engine crankshaft (11).
e. If desired, the spinner bulkhead (12) can be removed by removing screws and nuts attaching lugs
(13) to bulkhead. Note direction of lugs (13) and lug
attaching screws.
13-6. INSTALLATION.
a. If the spinner bulkhead (12) was removed, position bulkhead so the propeller blades will emerge
from the spinner (1) with ample clearance and install spinner bulkhead attaching lugs and screws.
CAUTION
Avoid scraping metal from bore of spinner
bulkhead and wedging scrapings between
engine flange and propeller. Trim the inside diameter of the bulkhead as necessary
when installing a new spinner bulkhead.
b. Clean propeller hub cavity and mating surfaces
of propeller and crankshaft.
c. Lightly lubricate a new O-ring (9) and the crankshaft pilot with clean engine oil and install the O-ring
in the propeller hub.

d. Align propeller mounting studs and dowel pins
with proper holes in engine crankshaft flange and
slide propeller carefully over crankshaft pilot until
mating surfaces of propeller and crankshaft flange
are approximately 1/4 inch apart.
e. Install propeller attaching nuts (14) and work
propeller aft as far as possible, then tighten nuts
evenly and torque to 660-780 lb-in.
f. Install any spacers (3) used between spinner
support and propeller cylinder, then install spinner
support and spinner. The spacers are used as required to cause a snug fit between the spinner (1)
and the spinner support (2).
13-7.

PROPELLER GOVERNORS.

13-8. DESCRIPTION. The propeller governor is a
single-acting, centrifugal type, which boosts oil pressure from the engine and directs it to the propeller
where the oil is used to increase blade pitch. A
single-acting governor uses oil pressure to effect a
pitch change in one direction only; a pitch change in
the opposite direction results from a combination of
centrifugal twisting moment of rotating blades and
compressed springs. Oil pressure is boosted in the
governor by a gear type oil pump. A pilot valve, fly
weight and speeder spring act together to open and
close governor oil passages as required to maintain
a constant engine speed.
NOTE
Outward physical appearance of specific
governors is the same, but internal parts
determine whether it uses oil pressure to
increase or decrease blade pitch. The
propellers used on these aircraft require
governors which "sense" in a certain manner. "Sensing" is determined by the type
pilot valve installed inside the governor.
Since the basic governor may be set to
"sense" oppositely, it is important to
ascertain that the governor is correct for
the propeller being used.
13-3

NOTE

,o
10

Use spacers (3) as required to
ensure a snug fit between spinner

Torque propeller mounting nuts
(14) to 660-780 lb-in.
3

14
13

With number 1 piston on top dead center,
position propeller with centerline of blades
vertical.

Detail C
DETAIL "C" APPLIES TO CYLINDER (4)
ATTACHMENT WHEN MODIFIED PER
SERVICE LETTER SE71-18

*THRU

AIRCRAFT SERIALS 18259421 AND A182-0116

BEGINNING WITH AIRCRAFT SERIALS 18259422 AND
A182-0117

Figure 13-1.
13-4

Propeller Installation

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

Spinner
Spacer
Cylinder
Screw
Propeller
Stud
Dowel Pin
O-Ring
Washer
Engine Crankshaft
Spinner Bulkhead
Lug
Mounting Nut
Screw
Tube
Safety Wire
Ring
Balance Weight
Balance Weight Bracket

Detail A

1.
2.
3.
4.
5.

Figure 13-2.

Propeller Governor
High-RPM Stop Screw
Governor Arm Extension
Nut
Control Rod End

6.
7.
8.
9.

Governor Control
Lever Assembly
Nut-Adjustment
Nut-Locknut

Governor and Control Adjustments

13-9. TROUBLE SHOOTING. When trouble shooting the propeller-governor combination, it is recommended that a governor known to be in good condition
be installed to check whether the propeller or the
governor is at fault. Removal and replacement, rigging, high-speed stop adjustment, desludging and replacement of the governor mounting gasket are not
major repairs and may be accomplished in the field.
Repairs to propeller governors are classed as propeller major repairs in Federal Aviation Regulations,
which also define who may accomplish such repairs.
13-10. REMOVAL,
a. Remove cowling and engine baffles as required
for access to governor.
b. Disconnect governor control from governor extension arm.
NOTE
Note EXACT position of all washers so that
washers may be installed in the same position on reinstallation.
c. Remove four sets of nuts and washers securing
governor to engine and pull governor from mounting
studs.
d. Remove gasket from between governor and engine mounting pad.

13-11. INSTALLATION.
a. Wipe governor and engine mounting pad clean.
b. Install a new gasket on the mounting studs. Install gasket with raised surface of the gasket screen
toward the governor.
c. Position governor on mounting studs, aligning
governor drive splines with splines in the engine and
install mounting nuts and washers. Do not force
spline engagement. Rotate engine crankshaft slightly
and splines will engage smoothly when properly
aligned.
d. Connect governor control to governor arm extension and rig control as outlined in paragraph 13-13.
e. Reinstall all items removed for access.
13-12. HIGH-RPM STOP ADJUSTMENT.
a. Remove engine cowling and baffles as required
for access.
b. Remove safety wire and loosen the high-speed
stop screw locknut.
c. Turn the stop screw IN to decrease maximum
rpm and OUT to increase maximum rpm. One full
turn of the stop screw causes a change of approximately 25 rpm.
d. Tighten stop screw locknut, safety wire stop
screw and make propeller control linkage adjustment
as necessary to maintain full travel.
e. Install baffles and cowling.
f. Test operate propeller and governor.

Change 1

13-5

NOTE
It is possible for either the propeller low
pitch (high-rpm) stop or the governor highrpm stop to be the high-rpm limiting factor.
It is desirable for the governor stop to limit
the high-rpm at the maximum rated rpm for
a particular aircraft. Due to climatic conditions, field elevation, low-pitch blade angle
and other considerations, an engine may not
reach rated rpm on the ground. It may be
necessary to readjust the governor stop after
test flying to obtain maximum rated rpm when
airborne.
13-13. RIGGING PROPELLER GOVERNOR CONTROL.
a. Disconnect governor control from governor
extension arm.
b. Place propeller governor control, in cabin,
full forward, then pull back approximately 1/8 inch
and lock in this position. This will allow "cushion"
to assure full contact of the governor arm with the
governor high-rpm stop screw.

13-6

c. Place governor arm against high-rpm stop
screw.
d. Loosen jam nuts and adjust control rod end
until attaching holes align while governor arm is
against high-rpm stop screw. Be sure to maintain
sufficient thread engagement of the control and rod
end. If necessary, shift control in the clamps to
achieve this.
e. Attach rod end to the governor arm extension.
Be sure all washers are installed correctly.
f. Operate the control to see that the governor arm
bottoms out against the low pitch stop and bottoms
out against or a maximum of .12 " from the high pitch
stop on the governor before reaching the end of control cable travel.
NOTE
The governors are equipped with an offset
extension to the governor arm. The offset
extension has an elongated slot to permit
further adjustment. The preceding steps
may still be used as an outline in the rigging
procedure.

SECTION 14
UTILITY SYSTEMS

TABLE OF CONTENTS
UTILITY SYSTEMS
............
Heating System
.
..........
Description ............
. ........
Operation ..
Trouble Shooting ..
.......
Removal, Installation and Repair..
Defroster System .
........
Description ............
Operation .............
Trouble Shooting ..........
Removal, Installation and Repair . . .
Ventilating System .
.........
Description ............
Operation .............
Trouble Shooting ..........
Removal, Installation and Repair. . .

14-1.

UTILITY SYSTEMS.

14-2.

HEATING SYSTEM.

Page
14-1
14-1
14-1
14-1
14-1
14-1
14-1
14-1
14-4A
14-4A
14-4A
14-4A
14-4A
14-4A
14-4A
14-4A

14-3. DESCRIPTION. The heating system is corprised of the heat exchange section of the exhaust
muffler, a shut-off valve, mounted on the right forward side of the firewall, a push-pull control on the
instrument panel, outlets and flexible ducting connecting the system.
14-4. OPERATION. Ram air is ducted through
engine baffle inlets and heat exchange section of the
exhaust muffler, to the shut-off valve at the firewall.
The heated air flows from the shut-off valve into a
duct across the aft side of the firewall, where it is
distributed into the cabin. The shut-off valve, operated by a push-pull control labeled "CABIN HEAT,"
located on the instrument panel, regulates the volume of heated air entering the system. Pulling the
control full out supplies maximum flow and pushing
control in gradually decreases flow, shutting off flow
completely when the control is pushed full in.
14-5. TROUBLE SHOOTING. Most of the operational troubles in the heating and defrosting systems are
caused by sticking or binding 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 must be repaired or replaced. When checking
controls, ensure valves respond freely to control
movement, that they move in the correct direction,
that they move through their full range of travel and

Oxygen System
............
14-5
Description ............
14-5
Maintenance Precautions ......
14-5
14-5
Replacement of Components .....
Oxygen Cylinder General Information .
14-6
Service Requirements
......
14-6
Inspection Requirements .....
14-6
Component Service Requirements .
14-6
Component Inspection Requirements .14-11
Masks and Hose ..
.
.
14-11
Maintenance and Cleaning .....
14-11
Purging ..............
14-11
Testing ..............
14-11
Leak Test .............
14-11
Charging ............
14-12

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 thorough
inspection of the exhaust system should be accomplished. Refer to applicable paragraph in Section 11
for this inspection. 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
Pro-Seal Co., Los Angeles, California) compound
or equivalent compound.
14-6. REMOVAL, INSTALLATION AND REPAIR.
Figure 14-1 may be used as a guide during removal,
installation and repair of heating system components.
Burned, frayed or crushed hose must be replaced
with new hose, cut to length and installed in the original routing. Trim hose winding shorter than the
hose to allow clamps to be fitted. Defective air
valves must be repaired or replaced. Check for
proper operation of valves and their controls after
repair or replacement.
14-7.

DEFROSTER SYSTEM.

14-8. DESCRIPTION. The defrosting system is
comprised of a duct across the aft side of the firewall, a defroster outlet and shut-off valve assembly
mounted on the left side of the cowl deck immediately
aft of the windshield, a shut-off valve control on the
instrument panel and flexible ducting connecting the
system.

Change 3

14-1

1. Cabin Heat Control
2. Nut
3. Washer
4. Arm
5. Roll Pin
6. Clamp Bolt
7. Spring

8.
9.
10.
11.
12.
13.
14.
15.

Figure 14-1.
14-2

Valve Plate Assembly
Valve Seat
Shim
Valve Body
Clamp
Hose
Screw
Deflector

Heating and Defrosting Systems

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

Cowl Deck
Nozzle
Cotter Pin
Valve
Shaft
Defroster Control
Duct

Nutplate
Air Scoop
Rib

16.
17.
18.

Insert
Fuselage Skin
Air Vent Door

30.
31.
32.

Directional Knob
Escutcheon
Valve Body

6.
7.
8.
9.
10.

Seal
Nut
Washer
Washer
Seal

20.
21.
22.
23.
24.

Seal
Inlet
Clamp
Hose
Setscrew

34.
35.
36.
37.
38.

Hose
Insulator
Spring
Cap
Seal

12.
13.
14.

Outlet Assembly
Knob
Washer

26.
27.
28.

Washer
Spring
Screw

40. Plate
41. Nut
42. Dome
43. Air Temperature Gage

2.
3.
4.

Detail A
OPTIONAL BEGINNING
WITH 18263366

Figure 14-2.

Ventilating Systems
Change 3

14-3

NOTE

Seal
22
Connector
Seal
Tube Assembly
(Inner)
48. Escutcheon
49. Housing
50. Valve Assembly
51. Wheel
NOTE
52. Headliner
Cessna Accessory Kit #AK182-191
53. Bracket
also installs the aft air vents.
54. Bracket
55. Retainer
44.
45.
46.
47.

Trap headliner (52) between housing
(49) and escutcheon (48). (Typical
entire perimeter of escutcheon.)
52
53

View

A-A

BEGINNING WITH SERIAL 18264296
Detail A

Figure 14-2.
14-4

Change 3

Ventilating Systems (Sheet 2 of 2)

14-9. OPERATION. Air from the duct across the
aft side of the firewall flows through a flexible duct
to the defroster outlet. The temperature and volume
of this air is controlled by the settings of the heater
system control.
14-10. TROUBLE SHOOTING. Since the defrosting
system depends on proper operation of the heating
system, refer to paragraph 14-5 for trouble shooting
the defrosting system.
14-11. REMOVAL, INSTALLATION AND REPAIR.
Figure 14-1 may be used as a guide during removal,
installation and repair of defrosting system components. Cut hose to length and install in the original routing. Trim hose winding shorter than the hose
to allow clamps to be fitted. A defective defroster
outlet must be repaired or replaced.
14-12.

VENTILATING SYSTEM.

14-13. DESCRIPTION. The ventilating system is
comprised of two airscoops mounted in the inboard
leading edge of each wing, a manually-adjustable
ventilator installed on each side of the cabin near the
upper corners of the windshield, two plenum chambers mounted in the rear cabin wing root areas, a
fresh airscoop door on the right side of the fuselage
just forward of the copilot's seat, a control knob on
the instrument panel and flexible ducting connecting
the system. Beginning with aircraft serial 18263366,
the outside air temperature gage may be located in
the right forward air vent. Refer to figure 14-2 for
removal and installation.
14-14. OPERATION. Air received from scoops
mounted in the inboard leading edges of the wing is
ducted to adjustable ventilators mounted on each side
of the cabin near the upper corners of the windshield,
Rear seat ventilation is provided by plenum chambers
mounted in the left and right rear cabin wing root
areas. These plenum chambers receive ram air from
the airscoops in the inboard leading edges of the

wings. Each plenum chamber is equipped with a
valve which meters the incoming cabin ventilation
air. This provides a chamber of expansion of cabin
air which greatly reduces inlet air noise. Filters
at the air inlets are primarily noise reduction filters.
Forward cabin ventilation is provided by a fresh airscoop door mounted on the right side of the fuselage,
just forward of the copilot seat. The scoop door is
operated by a control in the instrument panel marked
"CABIN AIR." Fresh air from the scoop door is
routed to the duct across the aft side of the firewall,
where it is distributed into the cabin. As long as
the "CABIN HEAT" control is pushed in, no heated
air can enter the firewall duct; therefore, when the
"CABIN AIR" control is pulled out, only fresh air
from the scoop will flow through the duct into the
cabin. As the "CABIN HEAT" control is gradually
pulled out, more and more heated air will blend with
the fresh air from the scoop and be distributed into
the cabin. Either one, or both of the controls may
be set at any position from full open to full closed.
14-15. TROUBLE SHOOTING. Most of the operational troubles in the ventilating system are caused
by sticking or binding of the inlet scoop door or its
control. Check the airscoop filter elements in the
wing leading edges for obstructions. The elements
may be removed and cleaned or replaced. Since air
passing through the filters is emitted into the cabin,
do not use a cleaning solution which would contaminate the air. The filters may be removed to increase air flow. However, their removal will cause
a slight increase in noise level.
14-16. REMOVAL, INSTALLATION AND REPAIR.
Figure 14-2 may be used as a guide during removal,
installation and repair of the ventilating system components. A defective ventilator or scoop must be repaired or replaced. Check for proper operation of
ventilating controls after installation or repair.

SHOP NOTES:

Change 3

14-4A/(14-4B (blank)

14-11.

OXYGEN SYSTEM.

WARNING
Under NO circumstances should the ON-OFF
control on the oxygen regulator be turned to
the "ON" position with the outlet (low pressure) ports open to atmosphere. Operation
of these units in this manner will induce
serious damage to the regulators and having
the following results:
1. Loss of outlet set pressure.
2. Loss of oxygen flow through the regulator which will result in inadequate oxygen being fed
through the aircraft system.
3. Internal leakage of oxygen through the
regulator.
Oepning of the control lever with the outlet ports
open to atmosphere, results in an "overshoot" of
the regulator metering device due to the extreme
flow demand through the regulator. After overshooting, the metering poppet device goes into oscillation,
creating serious damage to the poppet seat and diaphragm metering probe. This condition can occur
even by turning the control lever on and then turning
it quickly off.
A potential hazard exists to aircraft in the field where
inexperienced personnel might remove the cylinder
and regulator assembly from the aircraft and for
some reason, attempt to turn the regulator to the
"ON" position with the outlet ports open. Unfortunately, after the units have been improperly operated as
noted, there is no outward appearance indicating that
damage has occurred.
Testing these regulators should be accomplished only
after installation in the aircraft, with the "downstream" low pressure line attached.
14-12. DESCRIPTION. The system is comprised of
an oxygen cylinder and regulator assembly, filler
valve, pressure gage, pressure lines, outlets and
mask assemblies. The oxygen cylinder is mounted
aft of the baggage compartment. Locations of systemture
components are shown in figure 14-3. The pilot's
supply line is designed to receive a greater flow of
oxygen than the passengers. The pilot's mask is
equipped with a microphone, keyed by a switch button
on the pilot' s control wheel. The filler valve is located on the aft baggage curtain and access is gained
through the baggage door.

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.
14-13. 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 and parts of assemblies which
might, under any conditions, come in contact with
oxygen. The thread compound must be applied sparingly and carefully to only the first three threads of
the male fitting. No compound shall be used on aluminum flared fittings or on the coupling sleeves or
on the outside of the tube flares. The teflon tape
shall be used in accordance with the instructions
listed following this step. Extreme care must be
exercised to prevent contamination of the thread compound or teflon tape with oil, grease or other lubricants.
1. Lay tape on threads close to end of
fitting: Clockwise on standard threads,
opposite on left-hand 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. Resulted 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
Lines should be replaced by part
recommended
numbers called out in the aircraft arts Catalog
f Lies 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 obtained from American Mineral Spirits of Houston,
Texas.
N
Most air compressors are oil lubricated,
and a minute amount of oil may be carried
by the airstream. If only an oil lubricated
air compressor is available, drying must
be accomplished by heating at a temperaof 250° to 300°F 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, dry, 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 alkaline cleaner until free from oil and grease. Rinse with fresh water
and dry with clean, dry, filtered air.
NOTE
Cap lines at both ends immediately after
drying to prevent contamination.
14-14. REPLACEMENT OF COMPONENTS. Removal, disassembly, assembly and installation of
system components may be accomplished while using
figure 14-3 as a guide.
14-5

The pressure regulator, pressure gage and
line and filler valve should be removed and
replaced only by personnel familiar with
high-pressure fittings. Observe the maintenance precautions listed in the preceding
paragraph.
NOTE
Oxygen cylinder and regulator assemblies
may not always be installed in the field
exactly as illustrated in figure 14-3, 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 regulator. 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. Disassemble valve, replace O-rings and reassemble
valve. Install filler valve by reversing procedures
outlined in this step.
c. A cabin outlet is illustrated in figure 14-3. Repair kit, (part no. C166006-0108), available from
the Cessna Service Parts Center, may be used for
replacement of components of the outlet assembly.
d. To remove entire oxygen system, headliner
must be lowered and soundproofing removed to expose lines. Refer to Section 3 for headliner removal.
14-15. OXYGEN CYLINDER GENERAL INFORMATION. The following information is permanently
steel stamped on the shoulder, top head or neck of
each oxygen cylinder:
a. Cylinder specification, followed by service
pressure (e. g. "ICC-3AA1800" and "ICC-3HT1850"
for standard and light weight cylinders respectively).
NOTE
Effective 1 January 1970, all newly-manufactured cylinders are stamped "DOT"
(Department of Transportation), rather
than "ICC" (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
14-6

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
appears directly underneath the Cessna identification
placard.
14-16. 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
the date of the last hydrostatic test.
2. Light weight (ICC or DOT-3HT1850) cylinders must be hydrostatically tested to 5/3 their
working pressure every three years commencing
with the date of the last hydrostatic test.
b. Service life requirements:
1. Standard weight (ICC or DOT-3AA1800)
cylinders have no age life limitations and may continue to be used until they fail hydrostatic test.
2. Light weight (ICC or DOT-3HT1850) cylinders must be retired from service after 12 years
or 4, 380 filling cycles after date of manufacture,
whichever occurs first.
NOTE
These test periods and life limitations are
established by the Interstate Commerce
Commission Code of Federal Regulations,
Title 49, Chapter 1, Para. 73.34.
14-17. 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.
14-18. OXYGEN SYSTEM COMPONENT SERVICE
REQUIREMENTS.
a. PRESSURE REGULATOR. The regulator shall

MICROPHONE CABLE
Detail A
10

THROUGH SERIAL 18260055

TO FILLER VALVE

A
PILOT'S OXYGEN MASK

CABIN OUTLET

Detail
Detail

D
DetailB
Detail C

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

Base
Jam Nut
Spring
Poppet
Core
Escutcheon
Cover

8. Lock Ring
9. Low Pressure Relief Valve
10. Regulator
11. "ON-OFF" Control Cable
12. High Pressure Relief Valve
13. Pressure Gage
14. Seat
15. Piston

Figure 14-3.

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

O-Ring
Valve
Cap
Baggage Wall
Escutcheon
Cover
Bracket

Oxygen System (Sheet 1 of 5)
Change 3

14-7

SERIAL 18260056 THRU SERIAL 18260445

SEE SHEET 4

1. Filler Valve
2. Pressure Gage Line
3. "ON-OFF" Control Cable
4. Bulkhead Station 110.00
5. Oxygen Cylinder
7.
8.
9.

Regulator
Outlet
Overhead Console

SERIAL 18263476 THRU SERIAL 18264295

Figure 14-3.
14-8

Change 3

Oxygen System (Sheet 2 of 5)

SERIAL 18260446 THRU 18264295

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

Oxygen Cylinder
Fuselage Stringer
Bulkhead Station 140.00
Bracket
Bracket
Regulator

Figure 14-3.

Oxygen System (Sheet 3 of 5)
Change 3

14-9

SEE SHEET 1

SERIAL 18260056 THRU SERIAL 18264295

1.
2.
3.

Pressure Gage Line
Low Pressure Line
"ON-OFF" Control

4. Bracket
5. Cover
6. Speaker Grille
7. Arm

Figure 14-3.
14-10

Change 3

Oxygen System (Sheet 4 of 5)

8.
9.
10.

Knob
Outlet
Pressure Gage

14-3.
System (Sheet 5 of 5)ro
Oxygen
Figure
14-A/(14-B blank)
Change 3
7. Arm
8. Knob
10.

13. Support
14. oxygen Cylinder
15. Fuselage Stringer
16. Bulkhead Station 140.00
17. Regulator

Detail A

7
BEGINNING WITH SERIAL 18264296

Figure 14-3.

Pressure Gage

18.
19.

Tee
Filler Line

Oxygen System (Sheet 5 of 5)
Change 3

14-10A/(14-10B blank)

be functionally tested every two years or 1, 000 hours
for aircraft operating under 15, 000 ft. and one year
for aircraft operating over 15, 000 ft. The regulator
shall be overhauled every five years or at time of
hydrostatic test.
b. FILLER VALVE. The valve shall be functionally tested every two years and overhauled every five
years or at time of hydrostatic test.
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 checked
for accuracy and overhauled by an FAA approved
facility every five years.
e. OUTLETS. The outlets shall be disassembled
and inspected and the sealing core replaced, regardless of condition, every five years.
14-19. 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 build-up 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.
14-20. 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.
14-21. 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.

f. Replace hose if it shows evidence of deterioration.
g. Hose may be cleaned in the same manner as the
mask.
14-22. 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 14-25. 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.
14-23. 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 14-25.
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 75±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 75±10 PS.
d. Place oxygen control lever in OFF position and
allow test gage pressure to fall to 0 PSI. 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.
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 14-25.
14-24. 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

14-11

soap and water solution.
CAUTION
Do not attempt to tighten any connections
while the system is charged.
14-25.

SYSTEM CHARGING.

WARNINGBE 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.
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.

14-12

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 charged. Filling pressure refers to the pressure to which aircraft cylinders should be filled.
This table gives approximations only and assumes
a rise in temperature of approximately 25°F. due
to 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
psi pressure.

TABLE OF FILLING PRESSURES
Ambient
Temp.
°F
0
10
20
30
40

Filling
Press.
psig
1650
1700
1725
1775
1825

Ambient
Temp.
°F

Filling
Press.
psig

50
60
70
80
90

1875
1925
1975
2000
2050

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.
PRESSURE GAGE -

OXYGEN

OXYGEN PURIFIER
W/REPLACEABLE
CARTRIDGE

Figure 14-4.

Portable Oxygen Cascades

14-13 (14-14 blank)

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
SECTION 15
INSTRUMENTS AND INSTRUMENT SYSTEMS
TABLE OF CONTENTS

Page

INSTRUMENTS AND INSTRUMENT
SY ST EM S.................................................. 15-1
G eneral ................................................... 15-1
Instrument Panel .................................... 15-3
D escription......................................... 15-3
Removal and Installation ........................ 15-4
Shock-Mounts......................................... 15-4
Instruments............................................. 15-4
Removal.............................................. 15-4
Installation........................................... 15-4
Pitot and Static Systems ........................ 15-4
Description ........................................ 15-4
15-4
Maintenance.............................
Static Pressure System Inspection
and Leakage Test ............................ 15-4
Pitot System Inspection and
Leakage Test ................................... 15-7
Blowing Out Lines ............................... 15-7
Removal and Installation of
Components..................................... 15-7
Troubleshooting-Pitot-Static System... 15-8
True Airspeed Indicator....................... 15-8
Troubleshooting.............................. 15-8
Troubleshooting Altimeter ................... 15-10
Troubleshooting-Vertical Speed
Indicator ................................ ......... 15-10
Troubleshooting-Pitot Tube Heater ..... 15-11
Vacuum System ....................................... 15-11
Description ........................................ 15-11
Troubleshooting .................................. 15-11
Toubleshooting-Gyros ......................... 15-12
Troubleshooting-Vacuum pump .......... 15-15
Removal and Installation of
Components (Wet System).............. 15-15
Removal and Installation of
Components (Dry System) .............. 15-15
C leaning .............................................. 15-15
Vacuum Relief Valve Adjustment........ 15-15
Engine Indicators ............................................ 15-15
Tachometer ......................................... 15-15
Description ..................................... 15-15
Manifold Pressure Gage ..................... 15-16
15-1.

Description ........................................ 15-16
Troubleshooting ................................. 15-16
Cylinder Head Temperature Gage........ 15-17
Description ..................................... 15-17
Troubleshooting............................. 15-17
Oil Pressure Gage................................. 15-17
Description ...................................... 15-17
15-18
Troubleshooting .............................
Oil Temperature Gage .......................... 15-18
Description...................................... 15-18
Carburetor Air Temperature Gage........ 15-18
Description ...................................... 15-18
Troubleshooting .............................. 15-19
Fuel Quantity Indicating System........... 15-19
Description ..................................... 15-19
Removal and Installation of
Transmitter................................. 15-19
Troubleshooting .............................. 15-20
Transmitter Calibration ................... 15-20
15-20B
Hourmeter ........................................
.... 15-20B
Description ...........................
Economy Mixture Indicator ................... 15-20C
Description ...................................... 15-20C
Troubleshooting.............................. 15-21
..... 15-21
Calibration .............................
Removal and Installation ................ 15-22
Magnetic Compass ........................................ 15-22
D escription ............................................ 15-22
Stall Warning System and Transmitter ........... 15-22
D escription ............................................ 15-22
Turn Coordinator............................................. 15-22
..... ....................... 15-22
.
Description
. .....
....................... 15-22
Troubleshooting
Turn-and-Slip Indicator ................................... 15-23
Description ............................................ 15-23
Troubleshooting .................................... 15-23
........ ..................... 15-25
.
Electric Clock
............................ 15-25
.
Description
Wing Leveler ................................................. 15-25
.
............................ 15-25
Description
........ ..................... 15-25
.
Rigging

INSTRUMENTS AND INSTRUMENT SYSTEMS.

15-2. GENERAL. This section describes typical instrument installations and their respective operating systems.
Emphasis is placed on troubleshooting 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 Aviation Regulations 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
Revision 4
Mar 1/2004

15-1
© Cessna Aircraft Company

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
descriptive material, maintenance and troubleshooting information in this section is intended to help the mechanic
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 aircraft. Whether replacement is to be with a new instrument,
an exchange one, or original instrument is to be repaired must be decided on basis of individual circumstances.

15-1A/ (15-1B Blank)
© Cessna Aircraft Company

Revision 4
Mar 1/2004

BEGINNING WITH 18260446 AND
A18200137

THRU AIRCRAFT SERIAL 18261425 AND
BEGINNING WITH A18200137
17
16

Detail

A
Detail B

NOTE POSITION OF GROUND STRAP AND
SEQUENCE OF ATTACHING PARTS WHEN
REMOVING OR INSTALLING SHOCK PANEL

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

Marker Beacon Controls
Shock Mounted Panel
Removeable Panel
Radio and Switch Panel
Fuel and Engine Instruments
Knee Pad
Heating and Ventilating Controls
Wing Flap Control

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

Figure 15-1.
15-2

Change 1

NOTE
Detail A and B also
apply to sheet 2.

18260826 THRU 18261425 AND
BEGINNING WITH A18200137

Engine Controls
Circuit Breaker Panel
Switch Panel
Wing Leveler Control
Shock Mount
Ground Strap
Screw
Decorative Cover

Instrument Panel (Sheet 1 of 2)

Detail C

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

Panel
Spacer
Hook
Pile
Shim
Guide Pin
Rubber Grommet

BEGINNING WITH AIRCRAFT SERIAL 18261426

22
BEGINNING WITH
18261426

BEGINNING WITH
SERIAL 18262466
Detail

etail
Figure 15-1.

Instrument Panel (Sheet 2 of 2)

15-3. INSTRUMENT PANEL. (Refer to figure 15-1.)
15-4. 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 removeable panel,
containing flight instruments such as airspeed, verti-

Change 1

15-3

cal 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.
15-5. REMOVAL AND INSTALLATION.
a. FLIGHT INSTRUMENT PANEL.
1. (Thru 1971) Remove retainer clips securing
decorative cover by carefully prying under clip buttons. Do not drop spacers attached to clips. 1972
Models and on decorative covers are installed with
Velcro fasteners. 1974 models and on use a combination of Velcro fasteners and a pin and rubber
grommet arrangement to hold the decorative covers.
To remove pry loose and gently pull in a straight line.
2. Remove switch mounting nuts and switches
as necessary 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.

damage and entrance of foreign matter. Wire terminals should be insulated or tied up to prevent accidental grounding or short-circuiting.
15-9. 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 lubricant, shall be
installed using teflon tape on male fittings
only. This tape is available through the
Cessna Service Parts Center.
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 calibration of gages.
15-10. PITOT AND STATIC SYSTEMS.
figure 15-2.)

(Refer to

NOTE
Due to the difficulty encountered when removing 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.
1. Complete steps 1 and 2 above.
2. Tag and disconnect gyro plumbing.
3. Remove directional gyro mounting screws
and remove 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.
15-6. SHOCK-MOUNTS. Service life of shockmounted instruments is directly related to adequate
shock-mounting of the panel. If removel of shockmounted panel is necessary, check mounts for deterioration and replace as necessary.
15-7. INSTRUMENTS.
(Refer to figure 15-1.)
15-8. 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 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
15-4

Change 2

15-11. DESCRIPTION. The pitot system conveys
ram air pressure to the airspeed indicator. The
static system vents vertical speed indicator, altimeter and airspeed indicator to atmospheric pressure through plastic tubing connected to static ports.
A static line sump is installed at each source button
to collect condensation in static system. A pitot tube
heater may be installed. The heating element is controlled by a switch at the instrument panel and powered by the electrical system. A static pressure alternate source valve may be installed in the static systern for use when the external static source is malfunctioning. This valve also permits draining
condensate from the static lines. Refer to Owner's
Manual for flight operation using alternate static
source pressure. Beginning with 18263476, an encoding altimeter and a standby altimeter may be
installed. The encoding altimeter supplies an altimeter reading to the optional 300 or 400 transponder
for signal transmission. The standby altimeter is
connected to the static system by a tube to the vertical speed indicator. The static tube installation will
vary when an alternate static source is installed.
Figure 15-3A may be used as a guide for removal
and installation of the encoding altimeter system.
15-12. MAINTENANCE. Proper maintenance of
pitot and static system is essential for proper operation of altimeter, vertical speed and airspeed indicators. Leaks, moisture and obstructions in pitot
system will result in false airspeed indications,
while static system malfunctions will affect readings
of all three instruments. Under instrument flight
conditions, these instrument errors could be hazardous. Cleanliness and security are the principal
rules for system maintenance. The pitot tube and
static ports MUST be kept clean and unobstructed.

Use spacers (12) as required
for adequate friction on ring

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

Airspeed Indicator.
Altimeter
Vertical Speed Indicator
Static Line (To Right Sump)
Static Line (To Left Sump)
Pitot Line (To Pitot Tube)
Mounting Screw
Decorative Cover
Retainer
True Airspeed Ring
Instrument Panel
.
Spacer
Sump.
Static Port
Fuselage Skin
Heater Element (Heated Pitot Only)
Mast Body
Connector

.
7

.
..
.

...

.
-

6
18

TRUE AIRSPEED
INSTALLATION

-17
DetailB

Figure 15-2.

Pitot-Static Systems
15-5

THRU AIRCRAFT SERIALS
18260445 AND A182-0137
4

12.
13.

AIRCRAFT SERIALS 18260446
AND ON AND A182-0138 AND ON

*Valve (13) installed on some
1971 model aircraft.

Figure 15-3.
15-6

Alternate Static Air System

Valve
Valve

WITHOUT ALTERNATE STATIC SOURCE INSTALLED

WITH ALTERNATE STATIC SOURCE INSTALLED

BEGINNING WITH 18263476
1.
2.
3.
4.

Static Line
Standby Altimeter
Encoding Altimeter
Airspeed Indicator

6., Alternate Static Source
7. Line (To Transponder)

NOTE
NOTE

IF VERTICAL SPEED INDICATOR IS NOT
INSTALLED, ROUTE THE STATIC LINE

* TO R/H STATIC SOURCE
TO L/H STATIC SOURCE

Figure 15-3A

Encoding'Altimeter Installation

15-13. STATIC PRESSURE SYSTEM INSPECTION
AND LEAKAGE TEST. The following procedure
outlines inspection and testing of static pressure
system, assuming altimeter has been tested and inspected in accordance with current Federal Aviation

Regulations.
a. Ensure static system is free from entrapped
moisture and restrictions.
b. Ensure no alterations or deformations of airframe surface have been made which would affect

Change 2

15-6A/ (15-6B blank)

the relationship between air pressure in static pressure system and true ambient static air pressure for
any flight configuration.
c. Seal one static source port with pressure sensitive tape. This seal must be air tight.
d. Close static pressure alternate source valve,
if installed.
e. Attach a source of suction to the remaining static
pressure source opening. Figure 15-4 shows one
method of obtaining suction.
f. Slowly apply suction until altimeter indicates a
1000-foot increase in altitude.

When applying or releasing suction, do not
exceed range of vertical speed indicator or
airspeed indicator.
g. Cut off suction source to maintain a "closed"
system for one minute. Leakage shall not exceed
100 feet of altitude loss as indicated on altimeter.
lease suction source and remove tape from static
port.

15-14. PITOT SYSTEM INSPECTION AND LEAKAGE
TEST. To check pitot system for leaks, place a piece
of tape over small hole in lower aft end of pitot tube,
fasten a piece of rubber or plastic tubing over pitot
tube, close opposite end of tubing and slowly roll up
tube until airspeed indicator registers in cruise range.
Secure tube and after a few minutes recheck airspeed
indicator. Any leakage will have reduced the pressure in system, resulting in a lower airspeed indication. Slowly unroll tubing before removing it, so
pressure is reduced gradually. Otherwise instrument may be damaged. If test reveals a leak in sys15-15. BLOWING OUT LINES. Although the pitot
system is designed to drain down to pitot tube opening, condensation may collect at other points in system and produce a partial obstruction. To clear the
line, disconnect it at airspeed indicator. Using low
pressure air, blow from indicator end of line toward
the pitot tube.

Never blow through pitot or static lines toward
the instruments.

NOTE
If leakage rate exceeds the maximum allowable,
first tighten all connections, then repeat leakage test. If leakage rate still exceeds the maximum allowable, use following procedure.
i. Disconnect static pressure lines from airspeed
indicator and vertical speed indicator. Use suitable
fittings to connect lines together so altimeter is the
only instrument still connected into static pressure
system.
j. Repeat leakage test to check whether static pressure system or the bypassed instruments are cause of
leakage. If instruments are at fault, they must be
repaired by an "appropriately rated repair station"
or replaced. If static pressure system is at fault,
use following procedure to locate leakage.
k. Attach a source of positive pressure to static
source opening. Figure 15-4 shows one method of
obtaining positive pressure.

CAUTION
Do not apply positive pressure with airspeed
indicator or vertical speed indicator connected to static pressure system,
1. Slowly apply positive pressure until altimeter
indicates a 500-foot decrease in altitude and maintain this altimeter indication while checking for leaks.
Coat line connections and static source flange with
LEAK-TEC or a solution of mild soap and water,
watching for bubbles to locate leaks,
m. Tighten leaking connections. Repair or replace
parts found defective.
n. Reconnect airspeed and vertical speed indicators
into static pressure system and repeat leakage test
per steps "c" thru "h".

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 line to clear
with low pressure air.
NOTE
On aircraft equipped with alternate static
source, use the same procedure, opening
alternate static source valve momentarily
to clear line, then close valve and clear
remainder of system.
Check all static pressure line connections for tightness. If hose or hose connections are used, check
for general condition and clamps for security. Replace hose which have cracked, hardened or show
other signs of deterioration.
15-16. REMOVAL AND INSTALLATION OF COMPONENTS. (Refer to figure 15-2). To remove pitot
mast, remove four mounting screws on side of connector (18) and pull mast out of connector far enough
to disconnect pitot line (6). Electrical connections
to heater assembly (if installed) may be disconnected
through 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 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 nnd
right door. When replacing components of pitol and
static pressure systems, use anti-seize compound
sparingly on male threads on both metal and plaslie
connections. Avoid excess compound which might
enter lines. Tighten connections firmly. but avoid
overtightening and distorting fittings. If twistin of
15-7

15-17.

TROUBLE SHOOTING--PITOT-STATIC SYSTEM.
REMEDY

PROBABLE CAUSE

TROUBLE
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.

All three instruments - leaks
or obstruction in static line.

Repair or replace line.

Alternate static source valve
open.

Close for normal operation.

plastic tubing is encountered when tightening fittings,
VV-P-236 (USP Petrolatum), may be applied sparingly between tubing and fittings.
15-18. 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 15-2 for removal and installation. Upon in15-19.

TROUBLE SHOOTING.

stallation, before tightening mounting screws (7),
calibrate the instrument as follows: Rotate ring (10)
until 120 mph on adjustable ring aligns with 120 mph
on indicator. Holding this setting, move retainer (9)
until 60°F aligns with zero pressure altitude, then
tighten mounting screws (7) and replace decorative
cover.
NOTE
Beginning with aircraft serial 18264296, true
airspeed indicators are graduated in knots.
Therefore, use 105 knots instead of 120 miles
per hour in the above calibration procedure.

NOTE
Refer to paragraph 15-15 before blowing out pitot or
static lines.
TROUBLE
HAND FAILS TO RESPOND.

INCORRECT INDICATION OR
HAND OSCILLATES.

HAND VIBRATES.

15-8

Change 3

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.

Excessive vibration caused by
loose mounting screws.

Tighten mounting screws.

Excessive tubing vibration.

Tighten clamps and connections,
replace tubing with flexible hose.

NOTE
Air bulb with check valves may be obtained
locally from a surgical supply company. This
is the type used in measuring blood pressure.

PRESSURE

THICK-WALLED
SURGICAL HOSE

PRESSURE BLEED-OFF
SCREW (CLOSED)
AIR BULB
WITH CHECK-VALVES
CLAMP

SURGICAL HOSE-

-CHECK

SUCTION

CHECK VALVE

VALVE

TO APPLY SUCTION:
1.

Squeeze air bulb to expel as much air as possible.

Hold suction hose firmly against static pressure source opening.

Slowly release air bulb to obtain desired suction, then pinch hose shut tightly to trap suction in
system.

After leak test, release suction slowly by intermittently allowing a small amount of air to enter
static system. To do this, tilt end of suction hose away from opening, then immediately tilt it
back against opening. Wait until vertical speed indicator approaches zero, then repeat. Continue to admit this small amount of air intermittently until all suction is released, then remove
test equipment.

TO APPLY PRESSURE:

Do not apply positive pressure with airspeed indicator or vertical speed
indicator connected into static system.
1.

Hold pressure hose firmly against static pressure source opening.

Slowly squeeze air bulb to apply desired pressure to static system. Desired pressure may be
maintained by repeatedly squeezing bulb to replace any air escaping through leaks.

Release pressure by slowly opening pressure bleed-off screw, then remove test equipment.

Figure 15-4.

Static System Test Equipment
15-9

15-20.

TROUBLE SHOOTING -- ALTIMETER.
NOTE
Refer to paragraph 15-15 before blowing out pitot or
static lines.
TROUBLE

INSTRUMENT FAILS TO
OPERATE.

INCORRECT INDICATION.

HAND OSCILLATES.

15-21.

PROBABLE CAUSE

REMEDY

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
speed indicator installations.

Blow out lines, tighten connections.

TROUBLE SHOOTING -- VERTICAL SPEED INDICATOR.
NOTE
Refer to paragraph 15-15 before blowing out pitot or
static lines.
TROUBLE

INSTRUMENT FAILS TO
OPERATE.

INCORRECT INDICATION.

POINTER OSCILLATES.

15-10

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.

Partially plugged static line.

Blow out lines.

Leak in static line.

Repair or replace damaged lines,
tighten connections.

Leak in instrument case.

Replace instrument.

15-22.

TROUBLE SHOOTING -- PITOT TUBE HEATER.
OTE
Refer to paragraph 15-15 before blowing out pitot or
static lines.
TROUBLE

TUBE DOES NOT HEAT OR
CLEAR ICE.

15-23.

PROBABLE CAUSE
Switch turned "OFF. "

Turn switch "ON."

Popped circuit breaker.

Reset breaker.

Break in wiring.

Repair wiring.

Heating element burned out.

Replace element.

VACUUM SYSTEM (Refer to Figure 15-5.)

15-24. DESCRIPTION. Through Aircraft Serial
182060445 suction to operate the gyros is provided
by an engine-driven vacuum pump, gear-driven
through a spline-type coupling. The vacuum pump
discharge air passes through an oil separator, where
the oil, which passes through the pump for lubrication, is returned to the engine and the air is expelled
overboard. Beginning with Aircraft Serial 18260446
a dry vacuum system is installed. This system utilizes a sealed bearing, engine-driven vacuum pump,
which eliminates the oil separation components from
the system. A discharge tube is connected to the

15-25.

REMEDY

pump to expell the air from the pump overboard. A
suction relief valve is used to control system pressure and is connected between the pump inlet and the
instruments. In the cabin, the vacuum line is routed
from the gyro instruments to the relief valve at the
firewall. A central air filtering system is utilized.
Beginning with aircraft serial 18263970 a throw away
type central air filter will be installed using stastrap installation for quick change capability. The
reading of the suction gage indicates net difference
in suction before and after air passes through a gyro.
This differential pressure will gradually decrease as
the central air filter becomes dirty, causing a lower
reading on the suction gage.

TROUBLE SHOOTING -- VACUUM SYSTEM.
TROUBLE

PROBABLE CAUSE

REMEDY

HIGH SUCTION GAGE READINGS.
(Gyros function normally.)

Relief valve screen clogged,
relief valve malfunction.

Clean screen, 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.

Restriction in oil separator or
pump discharge line. (Wet system only. )

Clean oil separator.

Central air filter dirty.

Clean or replace filter.

Defective gage or sticking
relief valve.

Replace gage. Clean sticking valve
with Stoddard solvent. Blow dry
and test. If valve sticks after
cleaning, replace it.

SUCTION GAGE FLUCTUATES.

Change 1

15-11

15-26.

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.

DIAL SPINS IN ONE DIRECTION
CONTINUOUSLY.

15-12

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.

Operating limits have been
exceeded.

Replace instrument.

Defective mechanism.

Replace instrument.

Detail A

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

Gyro Horizon
Directional Gyro
Suction Gage
Central Air Filter
Hose (To Relief Valve)
Firewall
Suction Relief Valve
Hose (Oil Return)
Vacuum Pump
Oil Separator

WET VACUUM SYSTEM
THRU AIRCRAFT SERIAL 18260445

Figure 15-5.

Vacuum System (Sheet 1 of 2)
Change 1

15-13

13
Detail C
BEGINNING WITH
AIRCRAFT SERIAL
18263970

Detail D
BEGINNING WITH

AIRCRAFT SERIAL

118261894
1.

Gyro Horizon

6.
7.

Vacuum Pump
Overboard Drain Line

9. Tube Locator
10. Adapter Tube
11. Sta-Strap
12. Bracket
13. Connector
DRY VACUUM PUMP SYSTEM 14. Central Filter
BEGINNING WITH AIRCRAFT SERIAL 18260446
Figure 15-5.
15-14

Change 1

Vacuum System (Sheet 2 of 2)

15-27.

TROUBLE SHOOTING -- VACUUM PUMP.
REMEDY

PROBABLE CAUSE

TROUBLE

Damaged engine drive seal.

Replace gasket.

Oil separator clogged, oil
return line obstructed, excessive oil flow through pump.
(Wet system only)

Clean oil separator with Stoddard
solvent, then blow dry. Blow out
lines. If pump oil consumption is
excessive, replace oil metering
pin in pump.

HIGH SUCTION.

Suction relief valve
screen clogged.

Clean or replace screen.

LOW SUCTION.

Relief valve leaking.

Replace relief valve.

Vacuum pump failure.

Replace vacuum pump.

EXCESSIVE OIL IN DISCHARGE.

15-28. REMOVAL AND INSTALLATION OF COMPONENTS (WET SYSTEM). Through aircraft serial
18260445 the components of the vacuum system are
secured by conventional clamps, mounting screws
and nuts. To remove a component, remove mounting
screws and disconnect inlet and discharge lines,
When replacing a vacuum system component, ensure
connections are made correctly. Use thread lubricant sparingly and only on male threads. Avoid overtightening connections. Before reinstalling a vacuum
pump, probe oil passages in pump and engine, to
make sure they are open. Place mounting pad gasket
in position over studs and ensure it does not block oil
passages. Coat pump drive splines lightly with a
high-temperature grease such Dow Silicone #30
(Dow-Corning Co., Midland, Mich.). After installing pump, before connecting plumbing, start engine
and hold a piece of paper over pump discharge to
check for proper lubrication. Proper oil flow through
pump is one to four fluid ounces per hour.
15-28A. REMOVAL AND INSTALLATION OF COMPONENTS (DRY SYSTEM). Beginning with aircraft
serial 18260446 the components of the vacuum system
are secured by conventional clamps, mounting screws
and nuts. To remove a component, remove mounting
screws and disconnect inlet and discharge lines. Cap
open lines and fitting to prevent dirt from entering
the system. When replacing a vacuum system component, ensure connections are made correctly. Use
no lubricants on any components when assembling a
dry vacuum system. Avoid over-tightening connections. Before installing the vacuum pump, place
mounting pad gasket in position over studs. Be sure
all lines and fittings are open and caps are removed.

15-29.

CLEANING.

Low pressure, dry compressed

Never apply compressed air to lines or components installed in aircraft. The excessive
pressures will damage gyros. If an obstructed line is to be blown out, disconnect at both
ends and blow from instrument panel out.
15-30. VACUUM RELIEF VALVE ADJUSTMENT.
A suction gage reading of 5.3 inches of mercury is
desirable for gyro instruments. However, a range
of 4.6 to 5.4 inches of mercury is acceptable. To
adjust relief valve, remove central air filter, run
engine to 1900 rpm on ground and adjust relief valve
to 5.3 ± .1 inches of mercury.

Do not exceed maximum engine temperature.
Be sure filter element is clean before installing. If
reading drops noticeably, install new filter element.
15-31.

ENGINE INDICATORS.

15-32.

TACHOMETER.

15-33. 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 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 cable housing for

kinks, sharp bends and damage.

Disconnect cable

air should be used in cleaning vacuum system com-

at tachometer and pull it out of housing.

ponents.

for worn spots, breaks and kinks.

The suction relief valve should be washed

with Stoddard solvent then dried with low-pressure
air. Refer to Section 2 for central air filter. Check
hose for collapsed inner liners as well as external
damage.

Check cable

Change 1

15-15

15-34.

MANIFOLD PRESSURE GAGE.

15-35.

DESCRIPTION.

15-36.

TROUBLE SHOOTING.

is a barometric instrument which indicates absolute
pressure in the intake manifold in inches of mercury.

The manifold pressure gage

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.

Rocker shaft screws tight.

Replace instrument.

Link springs too tight.

Replace instrument.

Dirty pivot bearings.

Replace instrument.

Defective mechanism.

Replace instrument.

Leak in pressure line.

Repair or replace damaged
line, tighten connections.

Foreign matter in line.

Blow out line.

Damping needle dirty.

Replace instrument.

Leak in pressure line.

Repair or replace damaged line,
tighten connections.

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.

15-16

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
15-37.

CYLINDER HEAD TEMPERATURE GAGE

15-38. DESCRIPTION. The temperature sending unit regulates electrical power through the cylinder head
temperature gage. The gage and sending unit require little or no maintenance other than cleaning, making
sure lead is properly supported and all connections are clean, tight and properly insulated. When replacing a
sending unit, install as a matched pair. The 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 1520B when troubleshooting the cylinder head temperature gage.
NOTE:
15-39.

A Cylinder Head Temperature Gage Calibration Unit, (SK182-43) is available and may be
ordered through the Cessna Supply Division.

TROUBLESHOOTING
TROUBLE

PROBABLE CAUSE

REMEDY

GAGE INOPERATIVE

No current to circuit.
Defective gage, bulb or circuit.

GAGE FLUCTUATES RAPIDLY

Loose or broken wire permitting
alternate make and break of gage
circuit.

GAGE READS TOO HIGH ON
SCALE

High voltage.
Gage off calibration.

Check "A" terminal.
Replace gage.

GAGE READS TOO LOW ON
SCALE

Low voltage.
Gage off calibration.

Check voltage supply and "D"
terminal.
Replace gage.

GAGE READS OFF SCALE AT
HIGH END.

Break in bulb.
Break in bulb lead.
Internal break gage.

Replace bulb.
Replace bulb.
Replace gage.

OBVIOUSLY
READING

Defective gage mechanism.
Incorrect calibration.

Replace gage.
Calibrate system.

15-40.

INCORRECT

Repair electrical circuit.
Repair or replace defective
items.
Repair or replace defective
wire.

OIL PRESSURE GAGE

15-41.
DESCRIPTION. The Bourdon tube-type oil pressure gage is a direct-reading instrument, operated
by a pressure pickup line connected to the engine 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.

Revision 4
Mar 1/2004

15-17
© Cessna Aircraft Company

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
15-42. TROUBLESHOOTING.

TROUBLE

PROBABLE CAUSE

REMEDY

GAGE DOES NOT REGISTER

Pressure line clogged.
Pressure line broken.
Fractured Bourdon tube.
Gage pointer loose on staff.
Damaged gage movement.

Clean line.
Repair or replace damaged line.
Replace instrument.
Replace instrument.
Replace instrument.

Foreign matter in line.
Foreign matter in Bourdon tube.
Bourdon tube stretched.

Clean line.
Replace instrument.
Replace instrument.

Faulty mechanism.

Replace instrument.

Worn or bent movement.
Foreign matter in Bourdon tube.
Dirty or corroded movement.
Pointer bent and rubbing on dial,
dial screw or glass.
Leak in pressured line.

Replace
Replace
Replace
Replace

GAGE POINTER
RETURN TO ZERO

GAGE DOES
PROPERLY

NOT

GAGE
HAS
OPERATION

FAILS

ERRATIC

instrument.
instrument.
instrument.
instrument.

Repair or replace damaged line.

15-43. OIL TEMPERATURE GAGE.
15-44.
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 tubes inside diameter is small, small dents and kinks, which would be
acceptable in larger tubing, may partially or completely close off the capillary, making the gage inoperative.
Some airplanes are equipped with gages that are electrically actuated and are not adjustable. Refer to Table
1, on page 15-20A when troubleshooting the oil temperature gage.
15-45. CARBURETOR AIR TEMPERATURE GAGE.
15-46.
DESCRIPTION. The carburetor air temperature gage is of the resistance-bridge type. Changes in
electrical resistance of the element are indicated by the gage, calibrated for temperature. The system requires
power from the aircraft electrical system and operates only when the master switch is on. Although both
instrument and sensing bulb are grounded, two leads are used to avoid possibility of instrument error induced
by poor electrical bonds in the airframe.

15-18

Revision 4
© Cessna Aircraft Company

Mar 1/2004

15-47.

TROUBLE SHOOTING.
TROUBLE

GAGE POINTER STAYS OFF
LOW END OF SCALE.

PROBABLE CAUSE

REMEDY

Popped circuit breaker.

Reset breaker.

Master switch "OFF" or switch
defective.

Replace defective switch.

Broken or grounded leads
between gage and sensing
unit.

Repair or replace defective
wiring.

Defective gage or sensing unit.

Replace gage or sensing unit.

Broken or grounded lead.

Repair or replace defective wiring.

Defective gage or sensing unit.

Replace gage or sensing unit.

Defective master switch,
broken or grounded lead.

Replace switch, repair or
replace defective wiring.

Defective gage or sensing unit.

Replace gage or sensing unit.

Loose or broken lead

Repair or replace defective
wiring.

Defective gage or sensing unit.

Replace gage or sensing unit.

Excessive panel vibration.

Tighten panel mounting
screws.

OBVIOUSLY INCORRECT
TEMPERATURE READING.

Defective gage or sensing unit.

Replace gage or sensing unit.

POINTER FAILS TO GO OFF
SCALE WITH CURRENT OFF.

Defective master switch.

Replace switch.

Defective gage.

Replace gage.

GAGE POINTER GOES OFF
HIGH END OF SCALE.

GAGE OPERATES INTERMITTENTLY.

EXCESSIVE POINTER
OSCILLATION.

15-48.

FUEL QUANTITY INDICATING SYSTEM.

15-49. DESCRIPTION. The magnetic type fuel quanindicators are used in conjunction with a float tity
operated variable-resistance transmitter in each fuel
cell. The full position of float produced a minimum
resistance through the transmitter, permitting maximum current flow through the fuel quantity indicator
and maximum pointer deflection. As fuel level is
lowered, resistance in the transmitter is increased,
producing a decreased current flow through-the fuel
quantity indicator and a smaller pointer deflection.

15-49A. REMOVAL AND INSTALLATION TRANSMITTER. (Refer to section 12 figure 12-3).
a. Drain fuel from cell. (Observe the precautions
)
.......-- in Section 12, -paragraph 12-3.
b. Remove wing root fairing.
c. Disconnect electrical lead and ground strap from
transmitter.
d. Remove screws attaching transmitter and carefully work transmitter from cell. DO NOT BEND
FLOAT ARM.
e. Reverse preceding steps for installation. using
new gaskets around opening and under screw heads.

Change 1

15-19

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
15-50. TROUBLE SHOOTING.
TROUBLE
FAILURE TO INDICATE

PROBABLE CAUSE

REMEDY

No power to indicator or transmitter. (Pointer stays below E.)
Grounded wire. (Pointer stays
above F.)
Low voltage.
Defective indicator.

Check and reset breaker, repair or
replace defective wiring.
Repair or replace defective wire.

SYSTEM OFF CALIBRATION

Defective indicator.
Defective transmitter.
Low or high voltage

Replace indicator.
Recalibrate or replace.
Correct voltage.

STICKY OR SLUGGISH
INDICATOR OPERATION
ERRATIC READINGS

Defective indicator.
Low voltage
Loose or broken wiring on
indicator or transmitter.
Defective indicator or transmitter.
Defective master switch.

Replace indicator.
Correct voltage.
Repair or replace defective wire.

15-51.

Replace indicator or transmitter.
Replace switch.

TRANSMITTER ADJUSTMENT

WARNING:

15-51A.

Correct voltage.
Replace indicator.

USING THE FOLLOWING FUEL TRANSMITTER CALIBRATION PROCEDURES ON
COMPONENTS OTHER THAN THE ORIGINALLY INSTALLED (STEWART WARNER)
COMPONENTS WILL RESULT IN A FAULTY FUEL QUANTITY READING.

STEWART WARNER GAGE TRANSMITTER CALIBRATION
Chances of transmitter calibration changing in normal service is remote; however it is possible that
the float arm or the float arm stops may become bent if the transmitter is removed from the fuel
cell/tank. Transmitter calibration is obtained by adjusting float travel. Float travel is limited by the
float arm stops.

WARNING:

USE EXTREME CAUTION WHILE WORKING WITH ELECTRICAL COMPONENTS OF THE
FUEL SYSTEM. THE POSSIBILITY OF ELECTRICAL SPARKS AROUND AN "EMPTY"
FUEL CELL CREATES A HAZARDOUS SITUATION.

Before installing transmitter, attach electrical wires and place the master switch in the "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 the lower stop so pointer indicator is on E. Raise float
until arm is against upper stop to permit indicator pointer to be on F (full). Install transmitter in
accordance with paragraph 15-49A.
15-51B

ROCHESTER FUEL GAGE TRANSMITTER
Do not attempt to adjust float arm or stop. No adjustment is allowed.

15-20

Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL

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.

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.

Electrically ground the airplane.

Level the airplane and drain all fuel from wing fuel tanks.

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.

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
15-51A for instructions for adjusting Stewart Warner fuel indicating systems.
Rochester fuel quantity indicating system components are not adjustable, only
component replacement or standard electrical wiring system maintenance practices
are permitted.

With the fuel selector valve in the "OFF" position, add unusable fuel to each fuel tank.

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:

Revision 4
Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
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:

8.
15-51 D.

Stewart Warner fuel quantity indicating systems can be adjusted. Refer to paragraph 1515A for instructions for adjusting Stewart Warner fuel indicating systems. Rochester fuel
quantity indicating system components are not adjustable, only component replacement
or standard electrical wiring system maintenance practices are permitted.

Install any items and/or equipment removed to accomplish this procedure, remove maintenance
warning tags and connect the airplane battery.
OIL TEMPERATURE INDICATING SYSTEM RESISTANCE TABLE 1
The following table is provided to assist in troubleshooting the oil temperature indicating system
components.
Select the oil temperature sending unit part number that is used in your airplane from the left column
and the temperature from the column headings. Read the ohms value under the appropriate
temperature column.
Part Number
S1630-1
S1630-3
S1630-4
S1630-5
S2335-1

15-51E.

Type
Oil Temp
Oil Temp
Oil Temp
Oil Temp
Oil Temp

120°F

72°F

165°F

220°F
46.4

250°F
52.4
52.4

620.0
620.0
192.0

34.0

990.0

CYLINDER HEAD TEMPERATURE INDICATING SYSTEM RESISTANCE TABLE 2
The following table is provided to assist in the troubleshooting the oil temperature indicating system
components.
Select the cylinder head temperature sending unit part number that is used in your airplane from the
left column and the temperature from the column headings. Read the ohms value under the
appropriate temperature column.
Part Number
S1372-1
S1372-2
S1372-3
S1372-4
S2334-3
S2334-4

15-52.

Type
CHT
CHT
CHT
CHT
CHT
CHT

200°F

220°F
310.0
310.0

745.0
745.0

450°F
34.8
34.8
113.0
113.0

475°F
46.4

38.0
38.0

HOURMETER.

15-53.
DESCRIPTION. The hourmeter is an electrically operated instrument, actuated by a pressure
switch in the oil pressure gage line. Electrical power is supplied through a one-amp fuse from the electrical
clock circuit, and therefore will operate independent of the master switch.

Revision 4
Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
15-54. ECONOMY MIXTURE INDICATOR.
15-55. 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 entering the
engine cylinders. Refer to the Owner's Manual for operating procedure of the system.

Revision 4
Mar 1/2004

13.

BEGINNING WITH
1973

THRU 1972

Detail A
Figure 15-6.
15-56.

Metal Strip

Magnetic Compass Installation

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 15-57.

FLUCTUATING READING.

Loose, frayed or broken
lead, permitting alternate
make and break of current.

Tighten connections and
repair or replace defective
leads.

15-57. CALIBRATION. A potentiometer adjustment
screw is provided behind the plastic cap at the back
of the instrument for calibration. This adjustment
screw is used to position the pointer over the reference increment line (4/5 of scale) at peak EGT.
Establish 75% power in level flight, then carefully
lean mixture to peak EGT. After the pointer has peaked, using the adjustment screw, position the pointer

over reference increment line (4/5 of scale).

NOTE
This setting will provide relative temperature indications for normal cruise power
settings within range of the instrument.
Turning the screw clockwise increases the meter
reading and counterclockwise decreases the meter

reading.

There is a stop in each direction and damage

Change 1

15-21

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.
15-58. 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.
15-59.
15-6.)

MAGNETIC COMPASS.

(Refer to figure

15-60. 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 lamp. The compass mount is attached by three
screws to a base plate which is bonded to windshield
with methylene chloride. A tube containing the compass light wires is attached to the metal strip at the
top of the windshield. Removal of the compass is
15-65.

accomplished by removing the screw at forward end
of compass mount, unfastening the metal strip at the
top of windshield and cutting the two wire splices.
Removal of the compass mount is accomplished by
removing 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.
15-61.

STALL WARNING HORN AND TRANSMITTER.

15-62. DESCRIPTION. The stall warning horn is
mounted on the glove box. 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 16.
15-63.

TURN COORDINATOR.

15-64. 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 machanism 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.

Oil in indicator becomes
too thick.

Replace instrument.

Insufficient bearing end play.

Replace instrument.

Low voltage.

Correct voltage.

IN COLD TEMPERATURES,
HAND FAILS TO RESPOND
OR IS SLUGGISH.

15-22

15-65.

TROUBLE SHOOTING (Cont).

TROUBLE
NOISY GYRO.

15-66.

PROBABLE CAUSE
High voltage.

Correct voltage.

Loose or defective rotor
bearings.

Replace instrument.

TURN-AND-SLIP INDICATOR.

15-67. DESCRIPTION. The turn-and-slip indicator
is operated by the aircraft electrical system and

15-68.

REMEDY

operates ONLY when the master switch is on. Its
circuit is protected by an automatically-resetting
circuit breaker.

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.

15-23

NOTE

3 4

Inverter (2), turn coordinator (6)
and restrictor valve (5) must be
replaced as a matched set. For
field adjustment of restrictor valve
(5). refer to Brittian Level-Matic
Operation and Service Manual.

Torque hose mounting nuts
(13) to-12-14 Ib inches when

1. Hose (To Right Aileron)
2. Inverter
3. Hose (To Directional Gyro)
4. Hose (To Gyro Horizon)
5. Restrictor Valve
6. Turn Coordinator
7. Roll Trim Knob
8. ON-OFF Control Valve
9. Hose (To Left Aileron Servo)
10. Central Air Filter
11. Hose (To Relief Valve)
12. Bracket
13. Nut
14. Servo
15. Cable Guard

16
Detail

Change 1

B
17

20
.

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

Pulley
Clamp
Spring
Turnbuckle (Aileron
12
Direct Cable)
Bushing
Spacer
Bellcrank
THRU AIRCRAFT SERIAL 18262465
Bolt

Figure 15-7.
15-24

Wing Leveler Control System

15-68.

TROUBLE SHOOTING (Cont).

TROUBLE
NOISY GYRO.

15-69.

PROBABLE CAUSE

REMEDY

High voltage.

Correct voltage.

Loose or defective rotor
bearings.

Replace instrument.

ELECTRIC CLOCK.

15-70. 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.
15-71. WING LEVELER. (Refer to figure 15-7.)
(Through aircraft serial 18262465)
15-72. DESCRIPTION. A wing leveler control system, consisting of a turn coordinator (6), pneumatic
servos (14), connecting cables and hose (1 and 9) may
be installed. The turn coordinator gyro senses
changes in roll attitude, then electrically meters
vacuum power from the engine-driven vacuum pump
to the cylinder-piston servos, operating the ailerons
for lateral stability. Manual control of system is
afforded by the roll trim knob (7). The roll trim
should not be used to compensate for faulty rigging
or "wing heaviness". Manual override of the system
may be accomplished without damage to the aircraft

or system. The ON-OFF valve (8) controls vacuum
supply to the distributor valve, but does not affect
the electrically operated turn coordinator gyro. Installation of the wing leveler does not change the
vacuum relief valve settings. Refer to the appropriate publication issued by the manufacturer for trouble
shooting procedures.
15-73. RIGGING.
a. Remove access plates as necessary to expose
components.
b. Loosen clamp (17).
c. Move aileron to full UP position.
d. Move clamp (17) until outboard edge of clamp is
8.00 inches from center of bolt (23) and tighten clamp.
NOTE
After completon of step "d", servo seal
should be taut but not stretched.
e.

Repeat steps "a" through "d" for opposite wing.

SHOP NOTES:

Change 1

15-25/15-26(Blank)

SECTION 16
ELECTRICAL SYSTEMS
TABLE OF CONTENTS
ELECTRICAL SYSTEMS
General ...............

Page
.........

16-2
16-2

Electrical Power Supply System ....

16-2

Description ...........
Split Bus Bar ..........
Description .........
Split Bus Power Relay ......
Description .........
Master Switch ..........
Description .........
Ammeter
............
Description .........
Battery Power System ........
Battery ............
Description .........
Trouble Shooting .......
Removal and Installation . .
Cleaning the Battery .....
Adding Electrolyte or Water
to Battery .......
Testing the Battery .....
Charging the Battery .....
Battery Box ...........
Description .........
Removal and Installation . .
Maintenance of Battery Box . .
Battery Contactor. ........
Description .........

16-2
16-2
16-2
16-2
16-2
16-2
16-2
16-2
16-2
16-3
16-3
16-3
16-3
16-4
16-4

Removal and Installation

16-6

Battery Contactor Closing
Circuit ..........
Description .......
Ground Service Receptacle ......
Description ...........
Trouble Shooting .........
Removal and Installation .....
Alternator Power System .......
Description ...........
Alternator ............

16-4
16-4
16-6
16-6
16-6
16-6
16-6
16-6
16-6

16-19

Description

...........

16-9

Description ...........
Removal and Installation
Stall Warning System .........
Description

16-19
16-19
16-19
16-19
16-19
16-19
16-19
16-19
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30
16-30

(Beginning With 1973 Models)

16-12

...........

..........

Removal and Installation .....
Flashing Beacon .
..........
Description
.
..........
Removal and Installation ....
Anti-Collision Strobe Lights ......
Description
.
..........
Operational Requirements .....
Removal and Installation .....
Overhead Console .
.........
Description .
..........
Removal and Installation
.....
Instrument Lighting. .........
Description .
..........
Removal and Installation .....
Electroluminescent Panel Lighting . ..
Description
.
..........
Instrument Post Lighting
.......
Description .
..........
Removal and Installation .....
Transistorized Light Dimming .....
Description ...........
Removal and Installation .....
Dome Light .............
Removal and Installation .....
Map Light ..............

16-9

Removal and Installation

Description

16-19
16-19
16-19

16-6
16-6
16-8
16-8
16-8
16-9
16-9
16-9

16-10

. .

Description

....

Removal and Installation .....
Control Wheel Map Light
(Thru 1969 Models) .
......
Removal and Installation .....
Control Wheel Map Light
(1970 thru 1971 Models) ......
Description ...........
Removal and Installation .....
Control Wheel Map Light

Description .........
Trouble Shooting the Alternator
System
..........
Alternator Field Circuit
Protection
.........
Alternator Voltage Regulator .....
Description ...........
Trouble Shooting .........
Removal and Installation .....
Over-Voltage Warning System .....
Description ...........
Aircraft Lighting System .......
Description ...........
Trouble Shooting .........
Landing and Taxi Lights
(Thru 1971 Models) .......
Description ...........
Removal and Installation .....
Landing and Taxi Lights
(Beginning with 1972 Models .....

Removal and Installation .
Adjustment of Landing and
Taxi Lights ..........
Navigation Lights .
.......

...

.....

...........

16-30
16-30
16-30
16-33
16-33
16-33
16-33

16-33
16-33
16-33
16-33

16-33
16-33
16-33
16-33
16-33
16-36
16-36
16-36
16-36

16-19

Pitot and Stall Warning Heaters ...
Description ...........
Cigar Lighter .
.......
Description ...........
Removal and Installation .....
Emergency Locator Transmitter . .
Description
.
..........
Operation ............
Checkout Interval. ........
Removal and Installation of
Transmitter .
..
Removal and Installation of Antenna
Removal and Installation of
Magnesium Six Cell Battery Pack
Removal and Installation of
Lithium Four Cell Battery Pack
Trouble Shooting .
......

16-19

Electrical Load Analysis Chart

16-41

16-12
16-12
16-12
16-14
16-14
16-14
16-14
16-15
16-15
16-15
16-19
16-19
16-19

...

Change 3

16-38
16-38
16-38
16-39
16-39

16-1

16-1.

ELECTRICAL SYSTEMS.

16-2. GENERAL. This section contains service information necessary to maintain the Aircraft Electrical Power Supply System, Battery andExternal Power Supply System, Aircraft Lighting System, Pitot
Heater, Cigar Lighter and Electrical Load Analysis.
16-3.

16-8. DESCRIPTION. A power relay is installed
behind the instrument panel on all aircraft utilizing
a split bus bar. The relay is a normally closed type,
opening when external power is connected or when
the starter is engaged, thus removing battery power
from the electronic side of the split bus bar and preventing transient voltages from damaging the electronic installations. (See figure 16-1.)

ELECTRICAL POWER SUPPLY SYSTEM
16-9.

16-4. DESCRIPTION. Electrical energy for the aircraft is supplied by a 12-volt, direct-current, singlewire, negative ground electrical system. A single
12-volt battery supplies power for starting and furnishes a reserve source of power in the event of alternator failure. An engine-driven alternator is the
normal source of power during flight and maintains
a battery charge controlled by a voltage regulator.
An external power receptacle is offered as optional
equipment to supplement the battery system for starting and ground operation.
16-5.

SPLIT BUS BAR.

16-6. DESCRIPTION. Electrical power is supplied
through a split bus bar. One side of the bus bar supplies power to the electrical equipment while the
other side supplies the electronic installations. When
the master switch is closed the battery contactor engages and the battery power is supplied to the electrical side of the split bus bar. The electrical bus feeds
power to the electronic bus through a normally-closed
relay; this relay opens when the starter switch is engaged or when an external power source is used, preventing transient voltages from damaging the semiconductor circuitry in the electronics installations.
16-7.

SPLIT BUS POWER RELAY.

SHOP NOTES:

16-2

MASTER SWITCH.

16-10. DESCRIPTION. The operation of the battery
and alternator system is controlled by a master
switch. On models prior to 1970 the switch is a rocker type with double-pole, single-throw contacts. The
switch, when operated, connects the battery contactor
coil to ground and the alternator field circuit to the
battery, activating the power systems. On 1970
models and on, a new master switch is utilized.
This switch is an inter-locking 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" above the switch and is located on the left hand side of the switch panel.
16-11.

AMMETER.

16-12. 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 the battery is fully
charged and cruise is maintained with all electrical
equipment off, the ammeter will show a minimum
charging rate.

16-13. BATTERY POWER SYSTEM.
16-14. BATTERY.
16-15. DESCRIPTION. The battery is 12-volts and Is
approximately 33 ampere hour capacity. The battery is located in the tailcone and is equipped with
non-spill filler caps.
16-16.

TROUBLE SHOOTING
TROUBLE

BATTERY WILL NOT SUPPLY
POWER TO BUS OR IS INCAPABLE OF CRANKING ENGINE.

PROBABLE CAUSE

REMEDY

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 11.5 volts or
more. 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 battery at 20 amps
for approximately 30 minutes
or until the battery voltage
rises to 15 volts. Check battery with a load type tester.
If tester indicates a good battery, the malfunction may be
assumed to be a discharged
battery. If the tester indicates
a faulty battery, replace the
battery.

Faulty contactor or wiring
between contactor or 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.

Open coil on contactor.

4. Check continuity between
"BAT" terminal and master
switch terminal of contactor.
Normal indication is 16 to 24
ohms (Master switch open).
If ohmmeter indicates an open
coil, replace contactor. If
ohmmeter indicates a good
coil, proceed to step 5.

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 intermittant,
replace contactor. If voltage
is normal, proceed to step 6.
16-3

16-16.

TROUBLE SHOOTING (Cont).
PROBABLE CAUSE

TROUBLE

BATTERY WILL NOT SUPPLY
POWER TO BUS OR IS INCAPABLE OF CRANKING ENGINE
(cont).

Faulty wiring between contactor and bus.

16-17. REMOVAL AND INSTALLATION
(Refer to igure 16-1.)
a. Remove aft baggage wall.
b. Remove the battery box cover.
c. Disconnect the ground cable from the negative
battery terminal.
CAUTIoN
*When installing or removing battery always
observe the proper polarity with the aircraft electrical system (negative to ground).
Reversing the polarity, even momentarily,
may result in failure of semiconductor devices (alternator diodes, radio protection
diodes and radio transistors),
*Always remove the battery ground cable
first and replace it last to prevent accidental short circuits.
d. Disconnect the cable from the positive terminal
of the battery.
e. Lift the battery out of the battery box.
f. To replace the battery, reverse this procedure.CA
16-18. CLEANING THE BATTERY. For maximum
efficiency the battery and connections should be kept
clean at all times.
a. Remove the battery and connections in accordance with the preceding paragraph.
b. Tighten battery cell filler caps to prevent the
cleaning solution from entering the cells.
c. Wipe the battery cable ends, battery terminals
and the 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 with petroleum jelly

SHOP NOTES:

16-4

REMEDY

6. Inspect wiring between contactor and bus. Repair or replace wiring.

or an ignition spray product to reduce corrosion.
16-19. 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, however will decompose into gases and should be replaced regularly. Add distilled water as necessary to
maintain the electrolyte level with the horizontal baffle plate or the split ring on the filler neck inside the
battery. When "dry charged" batteries are put into
service fill as directed with electrolyte. When 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.
16-20. TESTING THE BATTERY. The specific
gravity of the battery may be measured with a hydrometer to determine the state of battery charge. If
the hydrometer reading is low, slow-charge the battery and retest. Hydrometer readings of the electrolyte must be compensated for the temperature of the
electrolyte. Some hydrometers have a built-in thermometer and conversion chart. The following chart
shows the battery condition for various hydrometer
readings with an electrolyte temperature of 80 °
Fahrenheit.

installed.
1. Split Bus Power Relay
2. Bracket - Relay Mounting
3. Screw
4. Washer
5. Spacer
6. Diode Board
7. Locknut
8. Nut
9. Lockwasher
10. Insulating Washer
Figure 16-1.

12.
13.
14.
15.
16.
17.
18.
19.
20.
21.

Fuse - Clock
Bracket - Fuse Mounting
Resistor
Diode
Solder Terminal
Battery
Battery Box Lid
Battery Box
Nylon Cover
Wire to Clock and Battery
Contactor Closing Circuit
Fuses

22. Diode Wire
23. Positive Battery Cable
24. Master Switch Wire
25. Bolt
26. Wire to Battery Contactor
Closing Circuit
27. External Power Cable
28. Battery Drain Tube
29. Clamp
30. Negative Ground Strap
31. Battery Contactor

Battery and Electrical Equipment Installation
Change 3

16-5

BATTERY HYDROMETER

READINGS
BATTERY
CONDITION

READINGS

1. 280 Specific Gravity ................

100% Charged

1. 250 Specific Gravity .................. 75% Charged

1. 190 Specific Gravity ..................

25% Charged

1. 160 Specific Gravity ............... Practically Dead

All readings shown are for an electrolyte
temperature of 80° Fahrenheit. For higher
temperatures the readings will be slightly
lower. For cooler temperatures the readings will be slightly higher. Some hydrometers will have a built-in temperature compensation chart and a thermometer. If this type
tester is used, disregard this chart.
16-21. CHARGING THE BATTERY. When the battery is to be charged, the level of the electrolyte
should be checked and adjusted by adding distilled
water to cover the tops of the internal battery plates,
Remove the battery from the aircraft and place in a
well ventilated area for charging.

WARNING
|

* When a battery is being charged, 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. Test the battery with a hydrometer to
determine the amount of charge. Decrease the
charging rate or stop charging temporarily if the
battery temperature exceeds 125°F.
16-22.

BATTERY BOX.

16-23. DESCRIPTION. The battery is completely
enclosed in an acid resistant plastic box which is
riveted to mounting brackets in the tailcone. The
box has a vent tube which protrudes through the bottom of the aircraft allowing battery gases and spilled
electrolyte to escape.
16-24. REMOVAL AND INSTALLATION.
(Refer to figure 16-1. ) The battery box is riveted to
16-6

Change 1

the mounting brackets in the tailcone. The rivets
must be drilled out to remove the box.
16-25. MAINTENANCE OF BATTERY BOX. 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.

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.

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
acid proof paint Part No. CES 1054-529, available
from the Cessna Service Parts Center.
16-26.

BATTERY CONTACTOR.

bolted to the side of the battery box. The contactor
is a plunger type contactor 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 transistorized radio equipment. The large terminal of the diode connects to the battery terminal of the
battery contactor. The small terminal of the diode
and the master switch wire connect to the coil terminal of the battery contactor. Nylon covers are installed on the contactor terminals to prevent accidental shorts. (See figure 16-1.)
16-28. REMOVAL AND INSTALLATION.
(Refer to figure 16-1.)
a. Remove the battery box cover and disconnect the
ground cable from the negative battery terminal and
pull cable clear of battery box.
b. Remove the nut, lockwasher and the two plain
washers securing the battery cables to the battery
contactor.
c. Remove the nut, lockwasher and the two plain
washers securing the wire which is routed to the master switch.
d. Remove the silicon diode which is connected to
the battery terminal and the coil terminal.
e. Remove the bolt, washer and nut securing each
side of the battery contactor to the battery box. The
contactor will now be free for removal.
f. To replace the contactor, reverse this procedure.
16-29.

BATTERY CONTACTOR CLOSING CIRCUIT.

16-30. DESCRIPTION. This circuit consists of a
5-amp fuse, a resistor and a diode mounted on a
bracket on the side of the battery box. This serves
to shunt a small charge around the battery contactor

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

Nipple
Lock Washer
Nut
Ground Strap
Washer
Bracket Assembly
Rivet
Doubler Assembly
Cowl
Door Assembly
Screw
Receptacle
Diode Board
Power Cable

Figure 16-2.

10
Detail A

Ground Service Receptacle Installation
16-7

so that ground power may be used to close the contactor when the battery is too dead to energize the
contactor by itself.

Adjust the supply for 14-volts and close the
master switch.
NOTE

16-31.

GROUND SERVICE RECEPTACLE.

16-32. DESCRIPTION. A ground service receptacle
is offered as optional equipment to permit 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

When using ground power to start the aircraft, close the master switch before removing the ground power plug. This will
ensure closure of the battery contactor
and excitation of the alternator field in the
event that the battery is completely dead.

is connected in series with the coil on the external
power contactor so that if the ground power source is
inadvertently connected with a reverse 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.

Failure to observe polarity when connecting
an external power source directly to the battery or directly to the battery side of the battery contactor, will damage the diodes in the
alternator and other semiconductor devices
in the aircraft.

IWARNING

NOTE
Maintenance of the electronic installation
cannot be performed when using external
power. Application of external power
opens the relay supplying voltage to the
electronic 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.

External power receptacle must be functionally
checked after wiring, or after replacement of
components of the external power or split bus
systems. Incorrect wiring or malfunctioned
components can cause immediate engagement
of starter when ground service plug is inserted.

16-33. TROUBLE SHOOTING.
TROUBLE

PROBABLE CAUSE

REMEDY

STARTER ENGAGES WHEN
GROUND POWER IS CONNECTED.

Shorted or reversed diode in
split bus-bar system.

Check wiring to, and condition
of diode mounted on the split
bus relay bracket adjacent to
the magneto switch. Correct
wiring. Replace diode board
assembly.

GROUND POWER WILL NOT
CRANK ENGINE.

Ground service connector
wired incorrectly.

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.

16-8

16-33.

TROUBLE SHOOTING. (Cont).
TROUBLE

GROUND POWER WILL NOT
CRANK ENGINE. (Cont).

PROBABLE CAUSE
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 16-24 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.

16-34. REMOVAL AND INSTALLATION.
(Refer to figure 16-2.)
a. Open the battery box and diconnect the ground
cable from the negative terminal of the battery and
pull the cable from the battery box.
b. Remove the nuts, washers, ground strap and
diode board from the studs of the receptacle and remove the battery cable.
c. Remove the screws and nuts holding the receptacle. The receptacle will then be free from the
bracket.
d. To install a ground service receptacle, reverse
this procedure. Be sure to place the ground strap
on the negative stud of the receptacle.
16-35.

REMEDY

ALTERNATOR POWER SYSTEM.

16-36. DESCRIPTION. The alternator system consists of an engine driven alternator, a voltage regulator mounted on the left hand side of the firewall 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. Beginning with 1972 models an over-voltage sensor
switch and red warning light labeled HIGH VOLTAGE
are incorporated to protect the system, (refer to paragraph 16-46). The aircraft battery supplies the source
of power for excitation of the alternator.

16-37.

ALTERNATOR.

16-38. DESCRIPTION. The 60-ampere alternators
used on the 182 model are three-phase, delta connected with integral silicon diode rectifiers. The
alternator is rated at 14-volts at 60-amperes continuous output. The moving center part of the alternator (rotor) consists of an axial winding with radial
interlocking poles which surround the winding. With
excitation applied to the winding through slip rings,
the pole pieces assume magnetic polarity. The rotor
is mounted in bearings and rotates inside the stator
which contains the windings in which the-ac is generated. The stator windings are three-phase, delta
connected, and are attached to two diode plates, each
of which contain three silicon diodes.
The diode plates are connected to accomplish fullwave, rectification of the ac. The resulting dc output is applied to the aircraft bus and sensed by the
voltage regulator. The regulator contorls the excitation applied to the alternator field, thus controlling
the output voltage of the alternator.

16-9

16-39.

TROUBLE SHOOTING THE ALTERNATOR SYSTEM.
TROUBLE

AMMETER INDICATES HEAVY
DISCHARGE WITH ENGINE
NOT RUNNING OR ALTERNATOR CIRCUIT BREAKER OPENS
WHEN MASTER SWITCH IS
TURNED ON.

PROBABLE CAUSE
Shorted radio noise filter
or shorted wire.

REMEDY
1. Remove cable from output
terminal of alternator. Check
resistance from end of cable
to ground (MASTER SWITCH
MUST BE OFF). If resistance
does not indicate a direct short,
proceed to step 4. If resistance
indicates a direct short, proceed
to step 2.
2. Remove cable connections
from radio noise filter. Check
resistance from the filter input
terminal to ground. Normal indication is infinite resistance.
If reading indicates a direct
short, replace filter. If no
short is evident, proceed to
step 3.
3. Check resistance from ground
to the free ends of the wires which
were connected to the radio noise
filter (or alternator if no noise
filter is installed). Normal indication does not show a direct short.
If a short exists in wires, repair
or replace wiring.

ALTERNATOR SYSTEM
WILL NOT KEEP BATTERY CHARGED.

16-10

Shorted diodes in alternator.

4. Check resistance from output
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.

Regulator faulty or improperly adjusted.

1. Start engine and adjust for
1500 RPM. Ammeter should
indicate a heavy charge rate
with all electrical equipment
turned off. Rate should taper
off in 1-3 minutes. A voltage
check at the bus should indicate
a reading consistant with the
voltage vs temperature chart
on page 16-14. 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 is low, proceed to step

16-39.

TROUBLE SHOOTING THE ALTERNATOR SYSTEM (Cont).

TROUBLE

ALTERNATOR SYSTEM
WILL NOT KEEP BATTERY CHARGED (Cont).

PROBABLE CAUSE

Regulator faulty or improperly adjusted. (Cont.)

REMEDY

2. Stop engine, remove cowl,
and remove cover from voltage
regulator. Turn master switch
ON/OFF several times and observe field relay in regulator.
Relay should open and close with
master switch and small arc
should be seen as contacts open.
If relay is inoperative, proceed
to step 3. If relay operates,
proceed to step 4.
3. Check voltage at "S" terminal
of regulator with master switch
closed. Meter should indicate bus
voltage. If voltage is present, replace regulator. If voltage is not
present, check wiring between
regulator and bus.
4. Remove plug from regulator
and start engine. Momentarily
jumper the "A+" and "F" terminals together on the plug. Ship's
ammeter should show heavy rate
of charge. If heavy charge rate
is observed, replace regulator.
If heavy charge rate is not observed, proceed to step 5.

Faulty wiring between alternator and regulator, or
faulty alternator.

5. Check resistance from "F"
terminal of regulator to "F" terminal of alternator. Normal
indication is a very low resistance. If reading indicates no,
or poor continuity, repair or
replace wiring from regulator
to alternator.
6. Check resistance from "F"
terminal of alternator to alternator case. Normal indication
is 6-7 ohms. If resistance is
high or low, repair or replace
alternator.
7. 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.

16-11

16-39.

TROUBLE SHOOTING THE ALTERNATOR SYSTEM (Cont.)
REMEDY

PROBABLE CAUSE

TROUBLE
ALTERNATOR OVERCHARGES
BATTERY - BATTERY USES
EXCESSIVE WATER.

Regulator faulty or improperly
adjusted.

Check bus voltage with engine
running. Normal indication
agrees with voltage vs temperature chart on page 16-13. Observe ship's ammeter, ammeter
should indicate near zero after a
few minutes of engine operation.
Replace regulator.

OVER-VOLTAGE WARNING
LIGHT ON.

Regulator faulty or improperly
adjusted. Faulty sensor switch.

1. With engine running turn off
and on battery portion of the
master switch. If the light stays
on shut down engine then turn on
the "BAT and "ALT" portions of
the master switch. Check for
voltage at the "S" terminal of the
voltage regulator. If voltage is
present adjust or replace regulator. If voltage is not present
check master switch and wiring
for short or open condition. If
wiring and switch are normal
replace sensor.

16-40. REMOVAL AND INSTALLATION.
(Refer to figure 16-3. )
a. Make sure the master switch remains in the off
position or disconnect the negative lead from the battery.
b. Disconnect and label the wiring from the alternator.
c. Remove the safety wire from the upper adjusting bolt and remove the bolt from the alternator.
d. Remove the nut and washer from the lower
mounting bolt.
e. Remove the alternator drive belt and the lower
mounting bolt to remove alternator.
f. To replace alternator, reverse this procedure.
g. Adjust belt tension to obtain 3/8" deflection at
center of belt when applying 12 pounds of pressure.
After belt is adjusted and the bolt is safety wired,
* tighten the bottom bolt to 100-140 lb-in. torque to remove any play between alternator mounting foot and
U-shaped support assembly. Whenever a new belt is
installed, belt tension should be checked within 10 to
25 hours of operation.
NOTE
When tightening the alternator belt, apply pry
bar pressure only to the end of the alternator
nearest the pulley.
16-41. ALTERNATOR FIELD CIRCUIT PROTECTION. On models prior to 1970, a 2-amp automatic
resetting circuit breaker located on the back of the
instrument panel is provided to protect the alternator
field circuit. On 1970 models and on, a manually16-12

Change 3

resettable circuit breaker located on the switch panel
is provided to protect the alternator field circuit.
16-42.

ALTERNATOR VOLTAGE REGULATOR.

16-43. DESCRIPTION. The alternator voltage regulator contains two relays. One relay is actuated by
the aircraft master switch and connects the regulator
to the battery. The second relay is a two-stage, voltage sensitive device, which is used to control the
current applied to the field winding of the alternator.
When the upper set of contacts on the voltage regulator relay are closed, full bus voltage is applied to the
field. This condition will exist when the battery is
being heavily charged or when a very heavy load is
applied to the system. When the upper contacts open,
as the voltage begins to rise toward normal bus voltage to the alternator field is reduced through a resistor network in the base of the regulator, thus reducing
the output from the alternator. As the voltage continues to rise, assuming a very light load on the system, the lower contacts will close and ground the alternator field and shut the alternator completely off.
Under lightly loaded conditions the voltage relay will
vibrate between the intermediate charge rate and the
lower (completely off) contacts. Under a moderate
load, the relay will vibrate between the intermediate
charge rate and the upper (full output) contacts.
The voltage relay is temperature compensated so that
the battery is supplied with the proper charging voltage for all operating temperatures. With the battery
fully charged (ship's ammeter indicating at or near

NOTE
At each 100 hour engine compartment inspection, Cessna Singleengine Service Letter SE71-42
dated December 10, 1971 should
be complied with.
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.

WARNING

Alternator
Adjusting Arm
Washer
Rubber Bushing
Bolt
Upper Adjusting Bolt
Drive Belt
Bushing
Bonded Mount
Nut
Support Assembly
Lower Mounting Bolt

On models manufactured prior to mid 1971
should alternator thru-bolt loosening or
breaking occur, Cessna Service Letter
SE71-40 dated November 24, 1971 should
be complied with. On models manufactured
after mid 1971 a new high strength thrubolt and a K shaped retainer are installed.
Torque bolts 45 to 55 pound-inches.

Figure 16-3.

Alternator Installation
Change 1

16-13

zero) and a moderate load applied to the system (a
taxi light turned on), the voltage at the bus bar should
be within the range shown according to the air temperature on the following chart:
Beginning with 18264296 a solid state voltage regulator is installed. The voltage limiter relay in this regulator is replaced by a circuit board. The regulator
is a remove and replace item and not repairable. The
regulator may be adjusted by removing the cover and
adjusting the potentiometer either up or down.
TEMPERATURE

BUS

13. 8 - 14. 1

75 - 90°F. ......

13.7 - 14.0

16-46.

OVER-VOLTAGE WARNING SYSTEM.

16-47. DESCRIPTION. Beginning with 1972 Models,
an over-voltage warning system is incorporated in the
aircraft. The over-voltage warning system consists
of an 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

60 - 74°F. ......

b. Remove the connector plug from the regulator.
c. Remove two screws holding the regulator on the
firewall.
d. To replace the regulator, reverse the procedure.
Be sure that the connections for grounding the alternator, wiring shields and the base of the regulator
are clean and bright before assembly. Otherwise,
poor voltage regulation and/or excessive radio noise
may result.

ment on the airplane is not recommended. A
bench
bench adjustment
adjustment procedure
procedure is
is outlined
outlined in
in the
the
Cessna Alternator Charging Systems Service/
Parts Manual.
The voltage 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.
16-44. TROUBLE SHOOTING. For trouble shooting
the voltage regulator, refer to paragraph 16-39.
16-45. REMOVAL AND INSTALLATION. (Refer to
figure 16-4. )
a. Make sure that the master switch is off, or disconnect the negative lead from the battery.

show a discharge. Turn off both sections of the MasSwitch to recycle the over-voltage sensor. If the
over-voltage condition was transient, the normal alternator charging will resume and no further action
is necessary. If the over-voltage tripout 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 can be tested by turning off the Alternator portion of the Master Switch and leaving the Battery
portion turned on. This test does not induce an overvoltage condition on the electrical system. On models
prior to aircraft serial 18260942, should nuisance
trip-outs occur caused by voltage spiks or transient
voltage, Cessna Single-engine Service Letter SE72-15
dated April 21, 1972 should be complied with.

4. screw

THRU 18264295
1. Voltage Regulator
2. Bolt
3. Firewall Shield

5.
6.
7.
8.
9.
10.
11.

Figure 16-4.
16-14

Change 3

Wire to Master Switch
Filter - Radio Noise
Wire to Filter
Wire to Alternator "F"
Wire to Alternator "A+"
Wire Shields to Ground
Wire to Alternator Ground

9
BEGINNI
WITH
18264296
12. Housing
13. Shields - Ground
14. Wire to Over-voltage Light

Voltage Regulator Installation

16-48.

AIRCRAFT LIGHTING SYSTEM.

16-49. DESCRIPTION. The aircraft lighting system
consists of landing and taxi lights, navigation lights,
flashing beacon light, anti-collision strobe lights,
dome and instrument flood lights, courtesy lights,
control wheel map light, compass and radio dial lights.
16-50.

On 1969 models & on, snap-in type rocker switches
are introduced. 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.

TROUBLE SHOOTING.
TROUBLE

LANDING AND TAXI LIGHTS
OUT.

LANDING OR TAXI LIGHT
OUT.

FLASHING BEACON DOES
NOT LIGHT.

FLASHING BEACON
CONSTANTLY LIT.

REMEDY

PROBABLE CAUSE
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 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.

Install new flasher.

Defective flasher.

Install new flasher.

Change 3

16-15

16-50.

TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

ALL NAV LIGHTS OUT.

ONE NAV LIGHT OUT.

ONE ANTI-COLLISION
STROBE LIGHT WILL
NOT LIGHT. THRU
1972 MODELS.

BOTH ANTI-COLLISION
STROBE LIGHTS WILL
NOT LIGHT. THRU
1972 MODELS.

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.

Open circuit in wiring.

2. Test wiring for continuity.
Repair or replace wiring.

Flash tube burned out.

Test with new flash tube.
flash tube.

Faulty wiring.

Test for continuity.
replace.

Faulty trigger head.

Test with new trigger head.
Replace trigger head.

Circuit breaker open.

Inspect.

Faulty power supply.

Listen for whine in power supply
to determine if power is operating.

Faulty switch.

Test for continuity.
replace.

Repair or

Faulty wiring.

Test for continuity.
replace.

Repair or

Inspect lamp.

Replace lamp.

Replace

Repair or

Reset.

WARNING
The 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.
BOTH ANTI-COLLISION
STROBE LIGHTS WILL
NOT LIGHT. BEGINNING
WITH 1973 MODELS.

16-16

Change 3

Open circuit breaker.

1. Check, if open reset. If
circuit breaker continues to
open proceed to step 2.

16-50.

TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

BOTH ANTI-COLLISION
STROBE LIGHTS WILL
NOT LIGHT. BEGINNING
WITH 1973 MODELS. Cont.

Open circuit breaker. Cont.

REMEDY
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.

(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
NOT LIGHT. BEGINNING
WITH 1973 MODELS.

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 to wing structure.
Check for 12 volts. If OK proceed
to step 2. If 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.

Change 3

16-17

16-50. TROUBLE SHOOTING (Cont).
TROUBLE
DOME LIGHT TROUBLE Cont.

INSTRUMENT LIGHTS
WILL NOT LIGHT.
(THRU 1969 MODELS).

PROBABLE CAUSE

REMEDY

Defective wiring Cont.

3. Test for open circuit. Repair
or replace wiring. If no short or
open circuit is found, proceed to
step 4.

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.

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.

SHOP NOTES:

16-18

Change 3

Defective rheostat.

4. Check voltage at instrument
light with master switch on.
Should read battery voltage with
rheostat turned full clockwise
and voltage should decrease as
rheostat is turned counterclockwise.
If no voltage is present or voltage
has a sudden drop before rheostat
has been turned full counterclockwise, replace rheostat.

Lamp burned out.

5. Test lamp with ohmmeter or
new lamp. Replace lamp.

16-50. TROUBLE SHOOTING (Cont).
TROUBLE

PROBABLE CAUSE

INSTRUMENT LIGHTS WILL
NOT LIGHT (1970 MODELS
& ON).

REMEDY

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.

INSTRUMENT LIGHTS WILL
NOT DIM (1970 MODELS &
ON.

CONTROL WHEEL MAP
LIGHT WILL NOT LIGHT
THRU 1969 AIRCRAFT
ONLY.

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.

Open resistor wiring in
minimum intensity end of
potentiometer.

1. Test for continuity. Replace
resistor or repair wiring.

Shorted transistor.

2. Test transistor by substitution.
Replces 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.

Inspect. Replace switch.

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.

5. Check voltage at map light
assembly with master and nav
switches on. If battery voltage
is present, replace map light
assembly.

CAUTION
Failure to observe polarity shown on wiring diagrams 11.11.0,
will result in immediate failure of the transistor on the map
light circuit board assembly.
Change 3

16-18A

16-50.

TROUBLE SHOOTING (Cont).
TROUBLE

CONTROL WHEEL MAP
LIGHT WILL NOT LIGHT
1970 AIRCRAFT & ON.

PROBABLE CAUSE

REMEDY

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.
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.

16-51. LANDING AND TAXI LIGHTS.
MODELS.)

(THRU 1971

16-52. DESCRIPTION. The landing and taxi lights
are mounted in the leading edge of the left wing. A
clear plastic cover provides weather protection for
the lamps and is shaped to maintain the leading edge
curvature of the wing. The landing lamp is mounted
on the inboard side and adjusted to throw its beam
further forward than the taxi light. Both lights are
controlled by a single switch.

SHOP NOTES:

16-18B

Change 3

5. Check voltage at map light
assembly with master and nav
switches on. If battery voltage
is present, replace map light
assembly.

16-53. REMOVAL AND INSTALLATION. (Refer to
figure 16-5. )
a. Remove the screws securing the landing light
window assembly and remove assembly.
b. Remove the four attaching screws (6) from the
bracket assembly and remove bracket.
NOTE
Do not reposition the landing and taxi light
adjustment screws (2). If adjustment is required, refer to figure 16-5.

c. Remove the two screws securing the wiring to
the lamp contacts and remove the lamp.
d. Install new lamp and reassemble.
16-54. LANDING AND TAXI LIGHTS. (BEGINNING
WITH 1972 MODELS.)
16-55. DESCRIPTION. Beginning with 1972 models
the landing and taxi lights are mounted in the lower
half of the engine cowl. Both lights are used for landing and the right hand for taxi. Lights are controlled
by a interlocking split rocker type switch thru 1973
models. Beginning with 1974 models two rocker type
switches are installed with a jumper wire and a diode
across the switches. With this arrangement the
switches operate the same as the interlocking split
rocker switch, the taxi light may be operated indivdually but when the landing lights are operated both landing and taxi lights are turned on.

16-61.

16-62. DESCRIPTION. The flashing beacon light is
attached to the vertical fin tip. The flashing beacon
is an iodine-vapor lamp electrically switched by a
solid-state flasher assembly. The flasher assembly
is located in the vertical fin under the fin tip. The
switching frequency of the flasher assembly operates
the lamp at approximately 45 flashes per minute.
On late 1970 models and on, a 1.5 ohm, 95 watt resistor has been added to the unused dual flasher lead
to provide a dummy load which is designed to eliminate a "pulsing" effect on the cabin lighting and ammeter.
16-63. REMOVAL AND INSTALLATION. For removal and installation of the flashing beacon light,
refer to figure 16-7.
16-64.

16-56. REMOVAL AND INSTALLATION. (Refer to
figure 16-5.)
a. Remove the lower cowl and disconnect wires
from the landing and taxi lamps.
b. Remove screws (8) securing lamp assembly to
support (2) and remove lamp assembly.
c. Remove screws (7) securing bracket (6) to
plate (3) and remove lamp.
d. To reinstall reverse this procedure.
16-57. ADJUSTMENT OF LANDING AND TAXI
LIGHT. Refer to figure 16-5. Adjustment of the
landing and taxi lights is pre-set at the factory
with adjustment screws bottomed out against the

bracket. Should Further adjustment be desired
proceed as follows.
a. Remove the lamp for access to adjustment
screws. (See figure 16-5. )
b. Thru 1971 Models adjustment is accomplished
by turning the screws until desired setting is obtained.
Beginning with 1972 Models washers must be added
on adjustment screws to change the setting.
NOTE
A maimum of two washers may be used
to adjust
setting.

Should removal of the cowling be desired to
make adjustments, ensure the landing and
taxi light wiring is disconnected before removing the bottom cowling.
c.

Remove cowling as outlined in Section 11.

16-58.

NAVIGATION LIGHTS.

FLASHING BEACON LIGHT.

ANTI-COLLISION STROBE LIGHTS.

16-65. DESCRIPTION. A white strobe light is installed on each wing tip. These lights are vibration
resistant and operate on the principle of a capacitor
discharge into a xenon tube, producing an extremely
high intensity flash. Energy is supplied to the strobe
lights from a power supply located inside the left wing
on the rib at wing station 136. 00, just aft of the landing light, thru 1972 models. Beginning with 1973
models each strobe light is equipped with its own
power supply, located on the wing tip ribs.
16-65A.

OPERATIONAL REQUIREMENTS.

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 plus, 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, 12 volt. Connect tube
to output of strobe power supply and allow to operate,
erating power supply at 12 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.
16-66. REMOVAL AND INSTALLATION. Refer to
figure 16-6 as a guide for removal and installation
of the anti-collision strobe light components.

WARNING
16-59. DESCRIPTION. The navigation lights are
located on each wing tip. The lights are controlled
by a single switch located on the instrument panel.
16-60. REMOVAL AND INSTALLATION. For removal and installation of navigation lights, refer to
figure 16-6.

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.

Change 3

16-19

D IM

E N S IO
NO.

182

1
2

0.68
0. 60

VIEW

A-A

2.
3.
4.
5.
6.

Figure 16-5.
16-20

Adjusting Screw
Lamp
Spring
Bracket
Screw

Landing and Taxi Light Installation (Sheet 1 of 2)

5
4.

Detail A
BEGINNING WITH 18261426

NOTE
A maximum of two washers
may be used between support
(2) and plate (3) for adjustment.

*
Detail A
18260826 THRU 18261425
1.
2.
3.
4.
5.
6.
7.
8.

Figure 16-5.

Nose Cap
Landing Light Support
Plate
Gasket
Lamp
Bracket
Tinnerman Screw
Screw

NOTE
A minimum of one gasket (4) and
a maximum of two gaskets may
be used to secure lamp (5) between bracket (6) and plate (3).

Landing and Taxi Light Installation (Sheet 2 of 2)
Change 1

16-21

12Da

Detail

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

Cap
Grounding Washer
Insulating Washer
Spring
Insulator
Lamp Socket
Gasket

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

Lamp
Lens
Lens Retainer
Screw
Clamp
Detector

Figure 16-6. Navigation and Anti-Collision Strobe Lights Installation (Sheet 1 of 2)
16-22

Change 1

* 18253594 THRU 182558505
* BEGINNING WITH 18258506

W.S. 136. 00

3
9

INSPECTION PLATE (REF)

LANDING
LIGHT (REF)

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

Electrical Leads
Cap
Washer
Insulated Washer
Spring
Insulator
Housing - Plug

Figure 16-6.

19
8.
9.
10.
11.
12.
13.
14.
15.

THRU 1972 MODELS
Housing- Cap
Wing Tip
Wing Navigation Light
Spacer
Flash Tube Assembly
Lens
Screw
Lens Retainer

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

Bulb
Seal
Bracket
Power Supply
Nutplate
Bolt
Wing Tip Rib
Gasket

Navigation and Anti-Collision Strobe Lights Installation (Sheet 2 of 2)
Change 1

16-23

late 1970 models and on.
15

12
1. Dome
3. Lamp
5. Baffle
6. Clamp Assembly

Detail A
iCAUTION_
When inserting lamp into socket
always use a handkerchief or a
tissue to prevent getting fingerprints on the lamp.

THRU 1972 MODELS
NOTE
*BEGINNING WITH 1973 MODELS
Fingerprints on lamp may shorten the life of the lamp.

Figure 16-7.
16-24

Change 1

Flashing Beacon Light Installation

8.
9.
10.
11.
12.
13.
14.
15.
16.

Nutplate
Tip Assembly
Spacer
Flasher
Fin Assembly
Housing - Cap
Housing - Plug
Plate
Dummy Load
(1. 5 Ohm Resistor)
17. Washer

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

DetailB

Adjusting Screw
Slide Knob
Cover Assembly
Grommet
Nut
Shield
Channel
Cover Plate

THRU 1969 MODELS ONLY.

Figure 16-8.

Overhead Console and Courtesy Light Installation (Sheet 1 of 2)
16-25

*THRU

1972 MODELS

*BEGINNING

2
WITH 1973 MODELS

.
1

Detail A
...

Detail

B
1970 MODELS AND ON.

Figure 16-8.
16-26

Screw

3. Nut
4.
5.

Housing Assembly
Grommet

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

Lamp
Cover Assembly
Oxygen Post Light
Slide Cover
Slide Knob
Shield
Cover Plate

Overhead Console and Courtesy Light Installation (Sheet 2 of 2)

THRU 1970 MODELS ONLY

1.
2.
3.

6.
7.
8.

Light Fitting Assembly
Nut
Light Assembly

Bracket
Gasket
Cover

Detail ABulb

5. Washer

Detail A

Figure 16-9.

Instrument Panel Glareshield Light Installation (Sheet 1 of 2)
16-27

BEGINNING WITH 1973 MODELS

1971 THRU 1972 MODELS
DETAIL

A TYPICAL FOR ALL POSITIONS

DETAIL

1.
2.
3.
4.
5.
Figure 16-9.
16-28

Nut
Lamp Assembly
Electrical Leads
Housing
Screw

Instrument Panel Glareshield Light Installation (Sheet 2 of 2)

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

Screw
Inverter
Washer
Nut
Glove Box
Transistor
Mica Washer
Heat Sink
Mounting Bracket
Housing-Socket

Figure 16-10.

Transistorized Light Dimming and Electroluminescent Light Inverter Installations

2
14

13
12

6.
7.
8.
10.
11.
12.

-13.

14.

Lens
Hood
Lamp-

Clear

Socket Assembly
Light Retainer
Switch
Tinnerman Nut
Adjustment Screw

THRU 1969 MODELS ONLY

Figure 16-11. Map Light Installation
16-29

16-67.

OVERHEAD CONSOLE.

16-68. DESCRIPTION. The overhead console contains a map light and the instrument flood lights.
The intensity of the instrument flood lights are controlled by a rheostat mounted on the switch panel.
The map light can be exposed by moving the slide
covers from the opening holes in the console.
16-69. REMOVAL AND INSTALLATION. For removal and installation, refer to figure 16-8.
16-70.

INSTRUMENT LIGHTING.

16-71. DESCRIPTION. The instrument panel lighting is fabricated in two separate sections. The
lower two-thirds of the instrument panel is illuminated by two lights mounted in the overhead light
console. The lighting for the upper one-third of the
instrument panel is provided by four (thru 1970 only)
or five (1971 and on) small lights located in the instrument panel glare shield. The intensity of the
instrument panel lighting is controlled by the instrument light dimming rheostat located on the switch
panel.
16-72. REMOVAL AND INSTALLATION. For removal and installation, refer to figure 16-8 and
16-9.
16-73. ELECTROLUMINESCENT PANEL LIGHTING.
16-74. 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 invertapak (power supply) located behind the instrument panel.
The intensity of the "EL" panel lighting is controlled
by a rheostat located on the instrument panel. (Refer
to 16-10).
16-75.

INSTRUMENT POST LIGHTING.

16-76. 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 are controlled by the radio light dimming rheostat located on
the switch panel.
16-77. REMOVAL AND INSTALLATION. For removal and installation of the post lamp, slide the cap
and lens assembly from the base. Slide the lamp
from the socket and replace.
16-78.

TRANSISTORIZED LIGHT DIMMING.

16-79. DESCRIPTION. A remotely located two-circuit, transistorized dimmer is installed as standard
equipment to control the instrument panel lighting on
1970 and on models. Panel lighting dimming controls
are increased from two to three. This is accomplished by concentric knob arrangement on one of the
existing control knobs. Transistor light dimming is
used on two of three circuits, thereby allowing greater dimming load variation and better linearity of
16-30

control. One circuit controls the engine instruments
and radio lights while the other circuit controls the
instrument flood lights and post lights.
16-80. REMOVAL AND INSTALLATION. For removal and installation, refer to figure 16-10.
16-81. DOME LIGHT. Thru 1972 models there are
two dome lights, one on each side of the cabin. Each
light assembly consists of a lens, lamp, socket and
retainer. Both dome lights are controlled by a single
switch located on the left rear door post. Beginning
with 1973 models the dome light is overhead just aft
of the console. The light is controlled by a switch on
the assembly.
16-82. REMOVAL AND INSTALLATION. Thru 1972
models for removal and replacement of dome lamps,
pry light assembly out of retainer then pry socket
out of light assembly. Twist the bayonet type lamp
from the socket and replace. Beginning with 1973
models the lens snap out for access to the lamp.
16-83. MAP LIGHTING.
16-84. DESCRIPTION. On models prior to 1970,
white map lighting and red, non-glare instrument
lighting are provided by an adjustable light mounted
on the side of the left forward door post. The switch
is a three-position type with red, white and off positions. The map light contains a white bulb for general purpose lighting and a red bulb for adjustable instrument lighting. The intensity of the red bulb is
controlled by the instrument light dimming rheostat
located on the switch panel. When instrument post
lights are installed, an extra map light mounted on
the right forward door post is included.
16-85. REMOVAL AND INSTALLATION. (Refer to
figure 16-11.)
a. For replacement of defective lamp, slide the
hood and lens from the map light assembly and remove the bayonet type lamp.
b. For removal of the map light assembly, remove
the screws from the front door post shield.
c. Remove the washer and nut attaching the map
light.
d. Remove the screw securing the ground wire.
e. Detach the wires at the quick disconnect fasteners
and remove the map light assembly.
f. To replace the map light assembly, reverse this
procedure.
NOTE
If map light swivels too freely, tighten the tension screw on the underside of map light.
16-86. CONTROL WHEEL MAP LIGHT. (THRU
1969 MODELS.) An optional control wheel map light
is available on the 1969 182 models. The map light
is mounted on the underside of the control wheel and
the light intensity is controlled by a thumb operated
rheostat. For dimming, the rheostat should be
turned clockwise.
16-87. REMOVAL AND INSTALLATION.
figure 16-12.)

(Refer to

Detail A

1. Socket - Lamp

13. Terminal Block

9. Rheostat
10. Control Wheel
11. Mike Key Switch

21. Resistor
22. Shield
23. Circuit Board

Detail B
Figure 16-12.

Control Wheel Map Light and Mike Key Switch Installation (Sheet 1 of 2)
16-31

* Plug (12) is used when mike

AIRCRAFT SERIAL 18260826

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

Cover
Adapter
Rubber Cover
Plate
Plate
Map
Map Light
Light Rheostat
Rheostat
Terminal Block
Map Light Assembly

Control Wheel

Figure 16-12.
16-32

Change 2

11.
12.
13.
14.
14.
15.
15.
16.
17.

Mike Switch
Plug
Insulator
Plug
Plug
Bracket
Bracket
Cable
Connector

18.

Knob (Map Light)

Control Wheel Map Light and Mike Key Switch Installation (Sheet 2 of 2)

a. For easy access of the map light assembly, rotate the control wheel 90 ° .
b. Remove the four screws from the map light circuit board. The map light assembly will then be free
for removal.
c. Label the wires connecting to the map light circuit board assembly and remove the screws securing
the wires to the circuit board assembly.
d. To install the map light reverse this procedure.
CAUTION
Failure to observe polarity shown on wiring
diagram (page 20-37), will result in immediate failure of the transistor on the map light
circuit board.
16-88. CONTROL WHEEL MAP LIGHT (1970 THRU
1971 MODELS.)
16-89. DESCRIPTION. Beginning with the 1970
models, a new type optional map light is installed
on the underside of the pilots control wheel. The
new map light consists of a rectangle shaped housing
containing two small lamps and a small rheostat.
16-90. REMOVAL AND INSTALLATION. (Refer to
figure 16-12.)
a. Rotate the control wheel 90 ° to the left to gain
access to the underside of the wheel.
b. Remove two screws and nuts holding map light
assembly to control wheel.
c. Detach two wires from the terminal strip above
the map light. Note the connection and mark for
reference when replacing the wires.
d. To install the control wheel map light reverse
this procedure.
e. For replacement of defective lamps, remove
two screws holding map light cover in place and
unplug rheostat to remove cover.
f. Unsnap lamp sockets and replace lamps.
g. To reassemble, reverse this procedure.

16-95. DESCRIPTION. The stall warning circuit is
comprised of a warning horn and an actuating switch.
The switch is installed in the leading edge of the left
wing and is actuated by airflow over the surface of
the wing. The switch will close as a stall condition
is approached, actuating the warning horn which is
mounted on the glove box. The stall warning unit
should actuate the stall warning horn approximately
five to ten miles per hour above the airplane stall
speed. Install the lip of the warning unit approximately one-sixteenth of an inch below the centerline
of the wing skin cutout. Test fly the aircraft to
determine if the unit actuates the warning horn at the
desired speed. If the unit actuates the warning horn
at a speed in excess of ten miles per hour above stall
speed, loosen the mounting screws and move the unit
down. If the unit actuates the horn five miles per
hour below stall speed, loosen the mounting screws
and move the unit up.
16-96.

PITOT AND STALL WARNING HEATERS.

16-97. DESCRIPTION. Electrical heater units are
incorporated in some pitot tubes and stall warning
switch units. The heaters offset the possibility of
ice formations on the pitot tube and stall warning
actuator switch. The heaters are integrally mounted
in the pitot tube and the stall warning actuator switch.
Both heaters are operated by the pitot heat switch.
16-98.

CIGAR LIGHTER.

16-99. DESCRIPTION. The cigar lighter (located
on the instrument panel) is equipped with a thermalactuated circuit breaker which is attached to the rear
of the cigar lighter. The circuit breaker will open if
the lighter becomes jammed in the socket or held in
position too long. The circuit breaker may be reset
by inserting a small probe into the . 078 diameter
hole in the back of the circuit breaker and pushing
lightly until a click is heard.
CAUTION

16-91. CONTROL WHEEL MAP LIGHT.
NING WITH 1972 MODELS.)

(BEGIN-

16-92. DESCRIPTION. The control wheel map light
is internally mounted in the control wheel. A rheostat switch located on the forward side of the control
wheel, thru 1974 models and on the lower side of the
control wheel beginning with 1975 models controls
the map light.
16-93. REMOVAL AND INSTALLATION. (Refer to
figure 16-12.) To remove, push upward on the lamp
and turn. The lamp and reflector is replaced as a
unit.
16-94.

Make sure the master switch is "OFF" before
inserting probe into the circuit breaker on
cigar lighter to reset.
16-100. REMOVAL AND INSTALLATION. (Refer to
figure 16-14. )
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.

STALL WARNING SYSTEM.

Change 2

16-33

BEGINNING WITH
1976 MODELS

DetailA
THRU 1975 MODELS
9

Detail B

~10.

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

Detail C

Figure 16-13.
16-34

Change 3

Pitot Heat and Stall Warning Installation

Glove Box
Screw
Nut
Washer
Bracket
Stall Warning Horn
Tinnerman Nut
Wing Leading Edge
Stall Warning Actuator
Heater Assembly
Pitot Tube

1. Knob
2. Element
3. Socket
4. Panel
5. Shell
6. Circuit Breaker
7. Probe
8. Nut
9. Lockwasher
10. Power Wire

Figure 16-14.

Cigar Lighter Installation

SHOP NOTES:

16-35

16-101. EMERGENCY LOCATOR TRANSMITTER.
16-102. DESCRIPTION. Several types of Emergency
Locator Transmitters (ELT) have been installed in
Cessna aircraft. Each of the ELT's is a self-contained, solid state unit, having its own power supply,
with an externally mounted antenna. The transmitters
are designed to transmit simultaneously on dual emergency frequencies of 121. 5 and 243. 0 Megahertz. All
units are 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 thru early 1974
models, were equipped with a battery-pack containing
six magnesium "D" size dry cell batteries wired in
series. (See figure 16-14) Mid 1974 thru early 1975,
ELT's are equipped with a battery-pack containing
four "in-line" lithium "D" size batteries wired in
series. Early 1975 and on ELT's are equipped with a
battery-pack containing four lithium "D" size batteries which are stacked in two's (See fig. 16-15).
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 MW-minimum), for
a continuous period of time as listed in the following
table.
TRANSMITTER LIFE
TO 75 MILLIWATTS OUTPUT

CAUTION
Do not leave the emergency locator transmitter
in the ON position longer than 5 seconds or
you may activate downed aircraft procedures
by C. A. P., D. 0. T. or F. A. A. personnel.

IWARNING
Magnesium (6-cell) battery-packs (excluding
4 cell lithium battery-packs) after prolonged
continuous use (1 hour) in a sealed environment give off explosive gas. If your ELT
has operated for this time period or longer,
as a precautionary measure, loosen the
ELT cover screws, lift the cover to break
air tight seal and let stand for 15 minutes
before tightening screws. Keep sparks,
flames and lighted cigarettes away from
battery-pack.
NOTE
After relatively short periods of inactivation,
the magnesium (6-cell) battery-pack develops
a coating over its anode which drastically
reduces self discharge and thereby gives
the cell an extremely long storage life.
This coating will exhibit a high resistance
to the flow of electric current when the
battery is first switched on. After a short
while (less than 15 seconds), the battery
current will completely dissolve this coating
and enable the battery to operate normally.
If this coating is present when your ELT is
activated, there may be a few seconds delay
before the transmitter reaches full power.
16-104. CHECKOUT INTERVAL:
100 HOURS.

Temperature

6 Cell
Magnesium
Battery Pack

+130°F
+ 70°F
- 4°F
- 40°F

89
95
49
23

hrs
hrs
hrs
hrs

4 Cell
Lithium
Battery Pack
115
115
95
70

hrs
hrs
hrs
hrs

Battery-packs have a normal shelf life of five to ten
(5-10) years and must be replaced at 1/2 of normal
shelf life in accordance with TSO-C91. Cessna
specifies 3 years replacement of magnesium (6-cell)
battery-packs and 5 years replacement of lithium
(4-cell) battery packs.
16-103. 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.
16-36

Change 3

a. Turn aircraft master switch ON.
b. Turn aircraft transceiver ON and set frequency
on receiver to 121. 5 MHz.
c. Remove the ELT's antenna cable from the ELT
unit.
d. Place the ELT's function selector switch in the
ON position for 5 seconds or less. Immediately replace the ELT function selector switch in the ARM
position after testing ELT.
e. Test should be conducted only within the time
period made up of the first five minutes after any
hour.
CAUTION
Tests with the antenna connected should be
approved and confirmed by the nearest control
tower.
NOTE
Without its antenna connected, the ELT will
produce sufficient signal to reach your receiver,
yet it will not disturb other communications
or damage output circuitry.

/cs^.='^i
*iL-*'.

-^^

^^-^

I|'"'^
INSTALLD AFT Of THIS PARTITION

Detail

Detail A

3. Fabric Fastener - Pile
4. Metal Strap
6. Bracket
7. Tailcone Skin
8. Sta-strap
9. Co-axial Cable

15.
16.
17.
18.

Detail A

Suppressor _W
Rubber Washer
Rubber Boot
Placard

Metal Strap (4) must be positioned so that
latch is on top of transmitter as installed
in the aircraft and not across transmitter
cover.

Figure 16-15.

Emergency Locator Transmitter Installation
Change 3

16-37

NOTE
After accumulated test or operation time
equals 1 hour, battery-pack replacement
is required.
f. Check calendar date for replacement of batterypack. This date is supplied on a sticker attached to
the outside of the ELT case and to each battery.
16-105. REMOVAL AND INSTALLATION OF TRANSMITTER. (Refer to figure 16-15.)
a. Remove the baggage curtain to gain access
to the transmitter and antenna.
b. Disconnect co-axial cable from end of transmitter.
c. Depending upon the particular installation, either
cut four sta-straps and remove transmitter or cut
sta-strap securing antenna cable and unlatch metal
strap to remove transmitter.
NOTE
Transmitter is also attached to the mounting
bracket by velcro strips; pull transmitter to
free from mounting bracket and velcro.

a. Disconnect co-axial cable from base of antenna.
b. Remove the nut and lockwasher attaching the
antenna base ot the fuselage and the antenna will be
free for removal.
c. To reinstall the antenna, reverse the preceding
steps.
NOTE
Upon reinstallation of antenna, cement
rubber boot (14) using RTV102, General
Electric Co. or equivalent, to antenna
whip only; do not apply adhesive to fuselage skin or damage to paint may result.
CAUTION
In-service 6 cell magnesium battery-pack
powered ELT's require the installation of a
static electricity suppressor in the antenna
cable to prevent the possibility of damage to
the case of the ELT. Refer to Cessna Avionics Service Letter AV74-16 and figure 16-13.
16-107. REMOVAL AND INSTALLATION OF MAGNESIUM SIX (6) CELL BATTERY-PACK. (Refer to
figure 16-16.)
NOTE

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, -do not allow solvent to dry on surface. Apply Velcro #40 adhesive to each surface in a thin even coat and allow to dry until
quite tacky, but no longer transfers to the
finger when touched (usually between 5 and
30 minutes). Porous surfaces may require
two coats. Place the two surfaces in contact
and press firmly together to insure intimate
contact. Allow 24 hours for complete cure.

On aircraft incorporating Cessna ELT's
manufactured by Leigh (Shark 7 series),
when replacing battery-pack refer to
Cessna Avionics Service Letter AV75-5
dated July 3, 1975.
NOTE
Since replacement 6 cell magnesium batterypacks are no longer available, when inservice units require replacement, use the
4 cell lithium battery-pack. Refer to paragraph 16-108.
TRANSMITTER
C589510-0102

e. To reinstall transmitter, reverse preceding
steps.
NOTE
An installation tool is required to properly
secure sta-straps on units installed with
sta-straps. This tool may be purchased
locally or ordered from the Pandiut Corporation, Tinley Park, Ill., part number
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.
16-106. REMOVAL AND INSTALLATION OF
ANTENNA. (Refer to figure 16-15.)
16-38

Change 3

ELECTRICAL
CONNECTOR

BATTERY-PACK
C589510-0105
(6 Cell Magnesium)

Figure 16-16. Magnesium 6 Cell
Battery-Pack Installation
16-108. REMOVAL AND INSTALLATION OF LITHIUM
FOUR (4) CELL BATTERY-PACK. (Refer to figure
16-17.)

NOTE
On aircraft incorporating Cessna ELT's
manufactured by Leigh (Shark 7 series)
when replacing battery-pack refer to
Cessna Avionics Service Letter AV75-5
dated July 3, 1975.

CAUTION
Be sure to enter the new battery-pack expiration date in the aircraft records. It is also
recommended this date be placed in your ELT
Owner's Manual for quick reference.

NOTE
Transmitters equipped with the 4 cell batterypack can only be replaced with another 4 cell
battery-pack.

TRANSMITTER
C589510-0202

BATTERY PACK
C589510-0205

a. After the transmitter has been removed from
aircraft in accordance with para. 16-105, 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 "O" ring gasket, rubber
washers and screws for reinstallation.

ELECTRICAL
CONNECTOR

JET MELT
ADHESIVE
3M (PN 3738)

c. Disconnect the battery-pack electrical connector
and remove battery pack.
d. Place new battery-pack in the transmitter with
four batteries as shown in the case in figure 16-17.
e. Connect the electrical connector as shown in
figure 16-17.
*NOTE
Before installing the new 4 cell batterypack, check to ensure that its voltage is
11. 2 volts or greater.
CAUTION
If it is desireable to replace adhesive material on the 4 cell battery-pack, use only 3M
Jet Melt Adhesive #3738. Do not use other
adhesive materials since other materials
may corrode the printed circuit board assembly.
f. Replace the transmitter cover by positioning the
rubber "O" ring gasket, if installed, on the cover
and pressing the cover and case together. Attach
cover with nine screws and rubber washers.
g. Remove the old battery-pack placard from the
end of transmitter and replace with new battery-pack
placard supplied with the new battery-pack.

TRANSMITTER
C589510-0209

BATTERY PACK
C589510-0210

Figure 16-17. Lithium 4 Cell
Battery Pack Installations

16-109. TROUBLE SHOOTING. Should your Emergency Locating Transmitter fail the 100 Hours performance checks, it is possible to a limited degree
to isolate the fault to a particular area of the equipment. In performing the following trouble shooting
procedures to test peak effective radiated power,
you will be able to determine if battery replacement
is necessary or if your unit should be returned to
your dealer for repair.

SHOP NOTES:

Change 3

16-39

TROUBLE
*POWER LOW

PROBABLE CAUSE

REMEDY

Low battery voltage.

1. Set toggle switch to off.
2. Remove plastic plug from the remote jack
and by means of a Switchcraft #750 jackplug,
connect a Simpson 260 model voltmeter and
measure voltage. If the battery-pack voltage
on the 6-cell magnesium battery pack transmitter is 10.8 volts or less, and on the 4-cell
lithium battery pack transmitters is 11.2 volts
or less, the battery pack is below specification.

Faulty transmitter.

3. If the battery-pack voltage meets the
specifications in step 2, the battery-pack
is 0. K. If the battery is O. K., check the
transmitter as follows:
a. Remove the voltmeter.
b. By means of a switchcraft 750 jackplug
and 3 inch maximum long leads, connect a
Simpson Model 1223 ammeter to the jack.
c. Set the toggle switch to ON and observe
the ammeter current drain. If the currentdrain is in the 85-100 ma range, the
transmitter or the co-axial cable is faulty.

Faulty co-axial
antenna cable.

4. Check co-axial antenna cable for high
resistance joints. If this is found to be
the case, the cable should be replaced.

*This test should be carried out with the co-axial cable provided with your unit.

SHOP NOTES:

16-40

Change 3

ELECTRICAL LOAD ANALYSIS CHART
STANDARD EQUIPMENT (Running Load)
Instrument Lights:
a. EL Panel .................
b. Cluster ..................
c. Console* .................
d. Compass .................
..................
e. Pedestal
....
....
..
Position Lights . ..
Battery Contactor ...............
Fuel Quantity Indicators ...........
Cylinder Head Temperature Indicators ..
Turn Coordinator# ...............

Clock . . . . . . . . . . . . . . . . . ..

1969

1970

AMPS REQD
1972 1973
1971

1974

1975

1976

. 75
0.3
2.0
0.1

.75
0.3
2.0
0.1
2
5.6
0.6
0. 4
0. 2
0. 8

.75
0.3
2.0
0. 1
.2
5.6
0. 6
0. 4
0. 2
0.. 8

.75
0.3
2.0
0.1
.2
5.6
0.6
0. 4
0. 2
0.88

.75
0.3
2.0
0.1
.2
5.6
0.6
0.4
0. 2
0.8

0.04
0.32
2.08
0.08
0.16
5.6
0.6
0.10
0. 05
0.8

0.04
0.32
2.08
0.08
0.16
5. 6
0. 6
0. 10
0. 05
0.8

0.04
0. 32
2.08
0.08
0.16
5.6
0.6
0.10
0. 05
0.8

6.5
0.03
-

6.5
4.0
0.03
-

6.5
4.0
0.03

1. 6

1.6

.02
4.5
4.5
--

1.0
1.0
.02

. 5.6
0.6
4
0. 2
8

...

OPTIONAL EQUIPMENT (Running Load)
Heated-Pitot (thru 1973) (Both Pitot and Stall
6.5
.....
........
warning 1974) ...
-Strobe Lights .
.
..............
0.03
Carburetor Air Temp .............
.Cessna 200A Navomatic (Type AF-295A)
Cessna 200A Navomatic (Type AF-295B) .....1. 6
Cessna 300 ADF (Type R-521B) .........
Cessna 300 ADF (Type R-546A) .........
Cessna 300 ADF (Type R-546E) ..............
02
Cessna 300 Marker Beacon (Type R-502B)
Cessna 300 Nav/Com (90 Channel-Type RT-517R) .4. 5
Cessna 300 Nav/Com (360 Channel-Type RT-540A). 4.5
Cessna 300 Nav/Com (100 Channel-Type RT-508A). Cessna 300 Nav/Com (360 Channel-Type RT-308C). -Cessna 300 Nav/Com (360 Channel-Type RT-528A).
Cessna 300 Nav/Com (360 Channel-Type RT-528E). -Cessna 300 Nav/Com (360 Channel-Type RT-328A). Cessna 300 Nav/Com (360 Channel-Type RT-328C .
Cessna 300 Nav/Com (720 Channel-Type RT-328D). 3.2
Cessna 300 Transceiver (Type RT-524A) .....
Cessna 300 HF Transceiver (Type PT10A). ....-5
Cessna 300 Transponder (Type KT-75R)...
Cessna 300 Transponder (Type KT-76 & KT-78) . ..
Cessna 300 Transponder (Type RT-359A) ....
2. 0
Cessna 300 Navomatic (Type AF-512C) ......
-Cessna 300 Navomatic (Type AF-512D) ......
Cessna 300 Navomatic (Type AF-394A) ......--..
Cessna 300A Navomatic (Type AF-395A) . ..
3.0
Cessna 300 DME (Type KN-60B) .........
-Cessna 300 DME (Type KN-60C) ........
2.0
. . . ..
Cessna 400 ADF (Type R-324A) ....
Cessna 400 ADF (Type R-346A) .........Cessna 400 ADF (Type R-446A) ........
. 5
Cessna 400 Glideslope (Type R-543B) ...
Cessna 400 Glideslope (Type R-443A) . .0.
Cessna 400 Glideslope (Type R-443B).
3.0
Cessna 400 Nav/Com (Type RT-522A). ......
Cessna 400 Nav/Com (Type RT-422A). ......Cessna 400 Transceiver (Type RT-532A) . . ..
1.5
Cessna 400 Transceiver (Type RT-432A) .....3. 0
Cessna 400 Transponder (Type RT-506A) . ....
Cessna 400 Transponder (Type RT-459A) .....-

6.5
4.0
0.03

1. 5
-2.0
--

3.2
1.5
1. 5
-

2. 0

2. 0
-

3.0
--

3.0

2.0

-3.0
-

1.0

0. 5

1.0
1.0
.02
1.9
-1. 9
1.9
1. 9--

1.0
1.0
.02
--1. 5

1. 5

-0. 5
-3.0
--1.5
-

3.0
-

2. 0
-1.0
1.0
.02
-

2.0
1.0
1.0
.02
-

1.9

0.5

3.2
1. 5
-1. 3
--

3.2
1.5
1.0

1.9

1. 5
3.2

1.5
3.2

3.0
-

3. 0
-1.5
3.0
-

1.0

2.0

-2.0
-3. 0

3.0

.
--

1.0
0.5

2.0

1.0
-1.3
0.5

4
3.0

10.0 10.0
4.0
4.0
0.03 0.03

1. 5
3.2
1.5
1.5
1.3

3.2

1. 9
1.9
-

10.0
4.0
0.03
2.0

3. 0
2.5
1. 5
1. 4
3.0
-

1.
0.5

0. 5

0. 4
3.0
2. 5
1.5
1.4
.0

0.4
3. 0

1. 0

Change 3

0.4
3.0

1.0
16-41

ELECTRICAL LOAD ANALYSIS CHART
OPTIONAL EQUIPMENT (Running Load) (CONT)
Sunair SSB Transceiver (Type ASB-125). ...
Flashing Beacon
..
....
King KN-60C DME
....
. . .. .
.
King KN-65 DME .
..........
..
Narco Mark 12A Nav/Com. ....
........
Narco Mark 12B Nav/Com with VOA-40 or VOA-51
Narco UGR-2 Glideslope Receiver ....
.
Pantronics PT10-A HF Transceiver ....
..
Brittain Wing Leveler ............
Post Lights* .................
Mkr Bcn E L Panel
.
. . . .........

IPS REQD
1972
1973

1969

1970

1971

7.0

5.0
7.0

4.6
4. 6
23

4. 6
.23

32
2.0

.32
2.0

1.52

10.0
.25

10.0
.25
1.0
3.3
15.6
15.0

10. 0
.25
1.0
3. 3
15.6
15. 0

10.0
.25
1.0
3. 3
15.6
15.0

10.0
.25
1.00
2. 5
15.6
15.0

10.0
.25
2.5
15.6
15.0

5.0
7.0

1974

1975

1976

5.0
7.0

5.0
7.0
3.0

5.0
7.0
2.8

1. 5
1.52

1.5
1.52
0.02

ITEMS NOT CONSIDERED AS PART OF
RUNNING LOAD
Cigarette Lighter ..
. ........
Stall Warning Horn ..............
Oil Dilution System .. ............
Wing Courtesy Lights and Cabin Lights . ....
Landing Lights . .............
Flap Motor .
. . ...........
tNegligible

16-42

.
.

3.3
15.6
15.0

*Only one or the other may be used at one time #Standard on Skylane Only

Change 3

10.0
.25

2. 5
15.6
15.0

2.5
15.6
15.0

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
SECTION 18
STRUCTURAL REPAIR
TABLE OF CONTENTS

Page

STRUCTURAL REPAIR

18-2

Repair Criteria ........................................

18-2

.18-2
Equipment and Tools ...................................
Support Stands ............................................. 18-2
18-2
Fuselage Repair Jigs ....................................
18-2
Wing Jigs ........................................
Wing Twist and Stabilizer AngleOf-Incidence ........................................ 18-2
Repair Materials .................................................. 18-2
W ing .................................................................... 18 -2
Description ........................................ 18-2
Wing Skin

18-2A

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

Negligible Damage .............................. 18-2A
Repairable Damage ............................. 18-2A
Damage Necessitating Replacement of Parts ................................
Wing Stringers ...............................
Negligible Damage ..............................
Repairable Damage .............................

18-2A
18-2A
18-2A
18-2A

Damage Necessitating Replacement of Parts ................................ 18-2A
Wing Auxiliary Spars .................................... 18-2A
Negligible Damage .............................. 18-2A
Repairable Damage ............................ 18-2A
Damage Necessitating Replacement of Parts.................................. 18-2A
18-2A
Wing Ribs ....................................
Negligible Damage ....................................... 18-3
Repairable Damage ..................................... 18-3
Damage Necessitating Replacement of Parts ........................................ 18-3
Wing Spars ...................................
18-3
Negligible Damage .............................. 18-3
Repairable Damage ............................. 18-3
Damage Necessitating Replacement of Parts ........................................ 18-3
Wing Leading Edge........................................18-3
Negligible Damage ...................................... 18-3
Repairable Damage ..................................... 18-3
Damage Necessitating Replacement of Parts ................................... 18-3
Bonded Leading Edges Repair ...................... 18-3
Negligible Damage. ............................ 18-3
Repairable Damage ............................ 18-3
A ilerons .......................................................... 18 -3
Negligible Damage .............................. 18-3
Cracks in Corrugated Aileron Skin ...... 18-3

Revision 4
Mar 1/2004

Repairable Damage ................................... 18-3A
Damage Necessitating Replacement of Parts ........................................ 18-3A
Aileron Balancing ....................................... 18-3A
Wing Flaps ........................................

Negligible Damage ..................................
Cracks in Corrugated Flap Skins................
Repairable Damage ...................................
Damage Necessitating Replacement of Parts ......................................

18-3A

18-3A
18-3A
18-3B
18-3B

Elevators and Rudder ........................................

18-3B

Negligible Damage ..................................

18-3B

CracKs in Corrugated Elevator

Skins ......... 18-3B

Repairable Damage ..............................
Damage Necessitating Replacement of Parts .....................................
Elevator and Rudder Balancing.............

18-4

Fin and Stabilizer .............................................

18-4

Negligible Damage................................

18-4

Repairable Damage ..............................

18-4

Damage Necessitating Replacement of Parts .....................................
......
Fuselage ......... ........................

18-4
18-4

D escription ................................................. 18-4

Negligible Damage................................
Repairable Damage ..............................
Damage Necessitating Replacement of Parts .....................................
Bulkheads .......................................................
Landing Gear Bulkheads.......................
Repair After Hard Landing ....................
Replacement of Hi-Shear Rivets......................
Firewall Damage...................................................

18-4
18-5

Engine M ount...................................................

18-5

18-5
18-5
18-5
18-5
18-5
18-5

18-5
Description ........................................
18-5
General Considerations ........................
18-5
Engine Mount Support Cradle Damage
ment of Parts .....................................
18-5
Damage Involving Engine Mounting
Lugs and Engine Mount to Fuselage
18-5
Attach Fittings....................................
B affles ................................................................... 18 -5
Engine Cowling ................................................
18-5
Repair of Cowling Skins............................. 18-5
Repair of Reinforcement Angles ................ 18-6
Repair of Thermo-Formed Plastic Components 18-6
Repair of Glass Fiber Constructed Components 18-6

18-1
©Cessn

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
18-1.

STRUCTURAL REPAIR.

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 wing attach 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.

WING TWIST AND STABILIZER ANGLE-OF-INCIDENCE.

18-8. Wing twist (washout) and horizontal stabilizer angle of incidence are shown below. Stabilizers do not
have twist. Wings have no twist from the root to the lift strut station. All twist in the wing panel occurs
between this station and the tip rib. Refer to figure 18-2 for wing twist measurement.
WING
Twist (Washout)
STABILIZER
Angle of Incidence

3°
-3° 30'

18-9. 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, 2024-T3 may be used in making repairs, since the strength of -T3 is greater
than -T4 or -T42 (-T4 and -T42 may be used interchangeably, but they may not be substituted for -T3). When
necessary to form a part with a smaller bend radius than the standard cold bending radius for 2024-T4, use
2024-0 and heat treat to 2024-T42 after forming. The repair material used in making a repair must equal the
gauge of the material being replaced unless otherwise noted. 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.
18-10. WING.
18-11
DESCRIPTION. The wing assemblies are a semicantilever type employing semimonocoque type of
structure. Basically, the internal structure consists of built-up front and rear spar assemblies, a formed
auxiliary spar assembly and formed sheet metal nose, intermediate, and trailing edge ribs. Stressed skin,
riveted to the rib and spar structures, completes the rigid 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

18-2

Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
access to aileron bellcranks, flap bellcranks, electrical wiring, strut attach fittings,control cables and pulleys,
and control disconnect points.
18-12. WING SKIN.
18-13. NEGLIGIBLE DAMAGE. Any smooth dents in the wing skin that are free from cracks, abrasions and
sharp corners, which are not stress wrinkles and do not interfere with any internal structure or mechanism,
may be considered as negligible damage. In areas of low stress intensity, cracks, deep scratches, or deep,
sharp dents, which after trimming or stop drilling can be enclosed by a two-inch circle, can be considered
negligible if the damaged area is at least one diameter of the enclosing circle away from all existing rivet
lines and material edges. Stop drilling is considered a temporary repair and a permanent repair must be
made as soon as practicable.
18-14. REPAIRABLE DAMAGE. Figure 18-4 outlines typical repair 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 de-burr. The sides of the hole should lie span-wise or chord-wise. circular patch may
also be used. If the patch is in an area where flush rivets are used, make a flush patch type of repair; if in an
area where flush rivets are not used, make an overlapping type of repair. Where optimum appearance and
airflow are desired, the flush patch may he used. Careful workmanship will eliminate gaps at butt-joints;
however, an epoxy type filler may be used at such joints.
18-15. 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 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.
18-16. WING STRINGERS.
18-17. NEGLIGIBLE DAMAGE. Referto paragraph 18-13.
18-18. 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.
18-19. DAMAGE NECESSITATING REPLACEMENT OF PARTS. If a stringer is so badly damaged that
more than one section must be spliced, replacement is recommended.
18-20. WING AUXILIARY SPARS.
18-21. NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.
18-22. REPAIRABLE DAMAGE. Figure 18-8 illustrates a typical auxiliary spar repair.
18-23. DAMAGE NECESSITATING REPLACEMENT OF PARTS. If damage to an auxiliary spar would
require a repair which could not be made between adjacent ribs, the auxiliary spar must be replaced.
18-24. WING RIBS.
18-25. NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.
18-26. REPAIRABLE DAMAGE. Figure 18-6 illustrates a typical wing rib repair.

Revision 4
Mmm DD/YYY

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
18-27.
DAMAGE NECESSITATING REPLACEMENT OF PARTS. Leading and trailing edge ribs that are
extensively damaged can be replaced. However, due to the necessity of unfastening an excessive amount of
skin in order to replace the rib, they should be repaired if practicable. Center ribs, between the front and rear
spar should always be repaired if practicable.
18-28.

WING SPARS.

18-29. NEGLIGIBLE DAMAGE. Due to the stress which wing spars encounter, very little damage can be
considered negligible. All cracks, stress wrinkles, deep scratches, and sharp dents must be repaired. Smooth
dents, light scratches and abrasions may be considered negligible.
18-30. REPAIRABLE DAMAGE. Figure 18-7, illustrates typical spar repairs. It is often practical to cut
repair pieces from service parts listed in the Parts Catalog. Service Kits are available for certain types of spar
repairs.
DAMAGE NECESSITATING REPLACEMENT OF PARTS. Damage so extensive that repair is not
18-31.
practicable requires replacement of a complete wing spar. Also refer to paragraph 18-2.
18-32.

WING LEADING EDGES.

18-33.

NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.

18-34.
REPAIRABLE DAMAGE. Wing skin repairs, outlined in paragraph 18-14, may be used to repair
leading edge skins, although the flush-type patches should be used. To facilitate repair, extra access holes
may be installed in locations noted in figure 18-13. If the damage would require a repair which could not be
made between adjacent ribs, refer to the following paragraph.
18-35.
DAMAGE NECESSITATING REPLACEMENT OF PARTS. Where extreme damage has occurred,
complete leading edge skin panels should be replaced. Extra access holes may be installed (refer to figure
18-13) to facilitate replacement.
18-36.

BONDED LEADING EDGES REPAIR.

18-37.

NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.

18-38. REPAIRABLE DAMAGE. (Refer to figure 18-11.) Cut out damaged area, as shown, to the edge of
undamaged ribs. Using a corresponding section from a new leading edge skin, overlap ribs and secure to
wing using rivet pattern as shown in the figure.
18-39.

AILERONS.

18-40.

NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.

18-40A. CRACKS IN CORRUGATED AILERON SKINS.
1.

It is permissible to stop drill crack(s) that originate at the trailing edge of the control surface
provided the crack is not more than 2 inches in length.

Stop drill crack using a #30 (0.128 inch) drill.

A crack may only be stop drilled once.
NOTE:

A crack that passes through a trailing edge rivet and does not extend to the trailing
edge of the skin may be stop-drilled at both ends of the crack.

18-3

Revision 4
Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
4.

Any control surface that has a crack that progresses past a stop-drilled hole shall be repaired.
Refer to paragraphs 18-40, 18-41, and 18-42 as applicable for repair information.

A control surface that has any of the following conditions shall have a repair made as soon as
practicable:
a.

A crack that is longer than 2 inches.

A crack that does not originate from the trailing edge of a trailing edge rivet.

Cracks in more than six trailing edge rivet locations per skin.
Refer to paragraphs 18-40, 18-41, and 18-42 as applicable for repair information.

Affected control surfaces with corrugated skins and having a stop drilled crack that does not
extend past the stop-drilled hole, may remain in service without additional repair.

18-41.
REPAIRABLE DAMAGE. The repair shown in figure 18-9 may be used to repair damage to aileron
leading edge skins. Figure 18-4 may be used as a guide to repair damage to the flat surface between
corrugations, when damaged area includes corrugations refer to figure 18-12. It is recommended that material
used for repair be cut from spare parts of the same gauge and corrugation spacing. Following repair the
aileron must be balanced. Refer to paragraph 18-43 for balancing. If damage would require a repair which
could not be made between adjacent ribs, refer to paragraph 18-42.
18-42.
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 preferable. Where extensive damage has occurred, replacement of
the aileron assembly is recommended. After repair and/or replacement, balance aileron in accordance with
paragraph 18-43 and figure 18-3.
18-43. AILERON BALANCING. Following repair, replacement or painting, the aileron must be balanced.
Complete instructions for fabricating balancing fixtures and mandrels and their use are given in figure 18-3.
18-44.

WING FLAPS.

18-45.

NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.

18-45A. CRACKS IN CORRUGATED FLAP SKINS.
1.

It is permissible to stop drill crack(s) that originate at the trailing edge of the control surface
provided the crack is not more than 2 inches in length.

Stop drill crack using a #30 (.128 inch) drill.

A crack may only be stop-drilled once.
NOTE:

A crack that passes through a trailing edge rivet and does not extend to the trailing
edge of the skin may be stop drilled at both ends of the crack.

Any control surface that has a crack that progresses past a stop-drilled hole shall be repaired.
Refer to paragraphs 18-45, 18-46, and 18-47 as applicable for repair information.

A control surface that has any of the following conditions shall have a repair made as soon as
practicable:
a.

A crack that is longer than 2 inches.

18-3A

Mar 1/2004

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
b. A crack that does not originate from the trailing edge of a trailing edge rivet.
c.

Cracks in more than six trailing edge rivet locations per skin.
Refer to paragraphs 18-45, 18-46, and 18-47 as applicable for repair information.

Affected control surfaces with corrugated skins and having a stop drilled crack that does not
extend past the stop drilled hole, may remain in service without additional repair.

18-46.

REPAIRABLE DAMAGE. Flap repairs should be similar to aileron repairs discussed in paragraph

18-41.

A flap leading edge repair is shown in figure 18-10.

18-47.
DAMAGE NECESSITATING REPLACEMENT OF PARTS. Flap repairs which require replacement
of parts should be similar to aileron repairs discussed in paragraph 18-42. Since the flap is not considered a
movable control surface, no balancing is required.
18-48.

ELEVATORS AND RUDDER.

18-49. NEGLIGIBLE DAMAGE. Refer to paragraph 18-13. 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.
18-49A. CRACKS IN CORRUGATED ELEVATOR SKINS.
1.

It is permissible to stop-drill crack(s) that originate at the trailing edge of the control surface
provided the crack is not more than 2 inches in length.

Stop-drill crack using a #30 (.0128 inch) drill.

A crack may only be stop-drilled once.
NOTE:

A crack that passes through a trailing edge rivet and does not extend to the trailing
edge of the skin may be stop-drilled at both ends of the crack.

Any control surface that has a crack that progresses past a stop-drilled hole shall be repaired.
Refer to paragraphs 18-45, 18-46, and 18-47 as applicable for repair information.

5. A control surface that has any of the following conditions shall have a repair made as soon as
practicable:
a.

A crack that is longer than 2 inches.

A crack that does not originate from the trailing edge of a trailing edge rivet.

Cracks in more than six trailing edge rivet locations per skin.
Refer to paragraphs 18-49, 18-50, and 18-51 as applicable for repair information.

Revision 4
Mar 1/2004

Affected control surfaces with corrugated skins and having a stop-drilled crack that does not
extend past the stop drilled hole, may remain in service without additional repair.

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
18-50.
REPAIRABLE DAMAGE. Skin patches illustrated in figure 18-4 may be used to repair skin damage
between corrugations. For skin damage which includes corrugations refer to figure 18-12. Following repair
the elevator/rudder must be balanced. Refer to figure 18-3 for balancing. If damage would require a repair
which could not be made between adjacent ribs, see paragraph 18-51.
18-51.
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 occurred,
replacement of the entire assembly is recommended. After repair and/or replacement, balance elevators and
rudder in accordance with paragraph 18-52 and figure 18-3.
18-52. ELEVATOR AND RUDDER BALANCING. Following repair, replacement or painting, the elevators
and rudder must be balanced. Complete instructions for fabricating balancing fixtures and mandrels and their
use are given in figure 18-3.
18-53.

FIN AND STABILIZER.

18-54.

NEGLIGIBLE DAMAGE. Refer to paragraph 18-13.

18-55.
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 on one side of the rear spar and ribs, and springing back the skin. If the damaged area would
require a repair which could not be made between adjacent ribs, or a repair would be located in an area with
compound curves, see the following paragraph.
18-56. 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.
18-57.

FUSELAGE.

18-58.
DESCRIPTION. The fuselage is of semimonocoque construction, consisting of formed bulkheads,
longitudinal stringer, reinforcing channels, and skin panels.
18-59.
NEGLIGIBLE DAMAGE. Refer to paragraph 18-13. 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 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 areas, wrinkles occurring 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 occurring 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.

Revision 4
Mar 1/2004

Wrinkles occurring in 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 1/18 inch of the nearest
structural members. Rivet pattern should be identical to existing manufactured seam at edge of sheet.
Negligible damage to stringers, formed skin flanges,
bulkhead channels, and like parts is similar to that
for the wing skin, given in paragraph 18-13.
18-60. REPAIRABLE DAMAGE. Fuselage skin repairs may be accomplished in the same manner as
wing skin repairs outlined in paragraph 18-14.
Stringers, formed skin flanges, bulkhead channels
and similar parts may be repaired as shown in figure 18-5.
18-61. DAMAGE NECESSITATING REPLACEMENT
OF PARTS. Fuselage skin major repairs may be
accomplished in the same manner as the wing repairs
outlined in paragraph 18-15. Damaged fittings must
be replaced. Seat rails serve as structural parts of
the fuselage and must be replaced if damaged.
18-62.

*Dash numbers to be determined according to the
size of the holes and the grip lengths required.
The bolts grip length should be chosen so that
no threads remain in the bearing area.
18-66. 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 stainless steel rivets.
Damaged or deformed angles and stiffeners may be
repaired as shown in figure 18-14, or they may be
replaced. A severely damaged firewall must be replaced as a unit.
18-67.

ENGINE MOUNT.

18-68. DESCRIPTION. The mount for the aircraft
engine is constructed of 4130 chrome-molybdenum
steel tubing. A truss structure, fastened to the firewall at four points, supports a cradle arrangement.
This cradle arrangement with its supporting lugs,
forms the base for rubber shock mounted engine supports.

BULKHEADS.

18-63. LANDING GEAR BULKHEADS. Since these
bulkheads are highly stressed members, irregularly
formed to provide clearance for control cables, fuel
lines, etc., the 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 damage must be repaired by
replacing the landing gear support assembly as an
aligned unit.
18-64. 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 examined, and all support
forgings must be checked for cracks, using a dye
penetrant and proper magnification. Bulkheads in
the damaged area must be checked for alignment,
and deformation of the bulkhead webs must be determined with the aid of a straightedge. Damaged support structure, buckled floorboards and skins, and
damaged or questionable forgings must be replaced.
18-65. REPLACEMENT OF HI-SHEAR RIVETS.
Hi-shear rivet replacement with close tolerance bolts
or other commercial fasteners of equivalent strength
properties is permissible. Holes must not be elongated, and the Hi shear substitute must be a smooth
push fit. Field replacement of main landing gear
forgings on bulkheads may be accomplished by using:
a. NAS464P* Bolt, MS21042-* Nut and AN960-*
washer in place of Hi-Shear Rivets for forgings with
machined flat surface around attachment holes.
b. NAS 464P* Bolt, ESNA 2935* Mating Base Ring,
ESNA LH 2935* Nut for forgings (with draft angle of
up to a maximum of 8 ° ) without machined flat surface
around attachment holes.

18-69. GENERAL CONSIDERATIONS., All welding
on the engine mount must be of the highest quality
since the tendency of vibration is to accentuate any
minor defect present and cause fatigure cracks. Engine mount members are preferably repaired by
using a larger diameter replacement tube, telescoped
over the stub of the original member using fishmouth
and rosette type welds. However, reinforced 30degree scarf welds in place of the fishmouth welds
are considered satisfactory for engine mount repair
work.
18-70. ENGINE MOUNT SUPPORT CRADLE DAMAGE. Minor damage such as a crack adjacent to an
engine attaching lug may be repaired by rewelding
the cradle tube and extending a gusset past the damaged area. Extensively damaged parts myst be replaced.
18-71. DAMAGE INVOLVING ENGINE MOUNTING
LUGS AND ENGINE MOUNT TO FUSELAGE ATTACHING FITTINGS. Engine mounting lugs and engine
mount-to-fuselage attaching fittings should not be repaired but must be replaced.
18-72. 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.
18-73.

ENGINE COWLING.

18-74. REPAIR OF COWLING SKINS. If extensively
damaged, complete sections of cowling must be replaced. Standard insert-type skin patches, however,
may be used if repair parts are formed to fit. Small
cracks may be stop-drilled and dents straightened if
they are reinforced on the inner side with a doubler
of the same material.
Change 2

18-5

1 X 12 X 30-3/4

X 12 X 48
X12 X 11
X12 X 8
30-3/4
2 X 4 X 20
5 INCH COTTON WEBBING--

----

3/8 INCH DIAMETER

BOLTS

2X4

NOTE
ALL DIMENSIONS ARE IN INCHES

Figure 18-1. Wing and Fuselage Support Stands
18-75. REPAIR OR REINFORCEMENT ANGLES.
Cowl reinforcement angles, if damaged, must be
replaced. Due to their small size they are easier
to replace than to repair.
18-76. REPAIR OF ABS COMPONENTS. Rezolin
Repair Kit, Number 404 may be obtained from the
Cessna Service Parts Center for repair of ABS
components.

18-6

Change 2

18-77. REPAIR OF GLASS-FIBER CONSTRUCTED
COMPONENTS. Glass-fiber constructed components
on the aircraft may be repaired as stipulated in instructions furnished in Service Kit SK182-12. Observe the resin manufacturer's recommendations
concerning mixing and application of the resing.
Epoxy resins are preferable for making repairs,
since epoxy compounds are usually more stable and
predictable than polyester and, in addition, give
better adhesion.

CESSNA AIRCRAFT COMPANY

MODEL 182 AND SKYLANE SERIES
SERVICE MANUAL
B3643

MEASURING WING TWIST
If damage has occurred to a wing, it is advisable to check the twist. The following method can be used with a
minimum of equipment, which includes a straightedge (32 inch minimum length of angle, or equivalent), three
modified bolts for a specific wing, and a protractor head with level.
1. Check chart for applicable dimension for bolt length (A or B).
2

Grind bolt to a rounded point as illustrated, checking length periodically.

Tape two bolts to straightedge according to dimension C.

4. Locate inboard wing station to be checked and make a pencil mark approximately one-half inch aft of the
lateral row of rivets in the wing leading edge spar flange.
5. Holding straightedge parallel to wing station (staying as clear as possible from "cans"), place longer bolt on
pencil mark and set protractor head against lower edge of straightedge.
6. Set bubble in level to center and lock protractor to hold this reading.
7. Omitting step 6, repeat procedure for each wing station, using dimensions specified in chart. Check to see
that protractor bubble is still centered.
8. Proper twist is present in wing if protractor readings are the same (parallel). Forward or aft bolt may be
lowered from wing 0.10 inch maximum to attain parallelism.
Figure 18-2. Checking Wing Twist
Revision 4
Mar 1/2004

BALANCING PROCEDURES
1.

Balance control surfaces in an enclosed draft free area.

Control surface to be balanced must be in the final flight configuration, painted (if applicable)
trim tabs installed, and all foreign matter removed from inside control surface.

If control surface is to be painted remove all existing paint prior to repainting and rebalancing.
Good workmanship and standard repair practices should not result in excessive additional
balance weight.

Place balancing mandrels (detail B) on a table or other suitable FLAT, LEVELED surface.
Mandrels must be placed at 90° to the hinge line of the control surface.

On control surfaces with the piano type hinges, insert inboard and outboard hinges into slotted
ends of the balancing mandrels, making sure that balancing mandrels are 90° to the hinge line.
On control surfaces with the bearing type hinge point, bolts or pins are inserted through the
attaching brackets, then placed on the knife edges of the mandrels as illustrated in (detail H).

AILERONS.
a.
(1) Block up the trailing edge of the aileron until a spirit-level protractor placed on the front
face of the aileron spar at W.S. 154.00 (± 6.00), (detail E), indicates 57° 10', (detail D).
(2)

ALTERNATE METHOD:
Measure the vertical distance from the aileron hinge point to the leveled surface.
1.80 inches, then block up trailing edge of the aileron to this measurement.

Subtract

b. With the aileron blocked in position place the balancing beam (detail A) at W. S. 154.00, (90° to
the hinge line), and adjust the trailing edge support on the balancing beam (detail D) until the
beam is level. If the aileron has not been disturbed during this operation, the beam is now
parallel to the aileron chord line at W. S. 154.00 (detail D).
NOTE
The above procedure must be performed with care. Small angular
discrepancies will produce large balancing errors.
c. Remove balancing beam and balance the beam by itself at the knife edges by adding washers
as shown, (detail C).
d.

Place the balancing beam on the aileron in its original position, then remove the blocks from
beneath the trailing edge.

e. Place the sliding weight (detail D) on the forward end of the balancing beam, moving it along
the beam until the beam is again level. A small, lightweight, spirit-level may be used for
this purpose provided it is symmetrical about its bubble reference and this reference is
placed on the beam directly over the aileron hinge line (detail D).
f. If aileron is correctly balanced, the position of the sliding weight with respect to the aileron
hinge line, will produce a moment about the hinge line somewhere within the underbalance
tolerance listed in the chart on (Sheet 5 of 5).
g. If modification of the aileron balance weight is necessary to correct an out-of-tolerance
condition, the balance weight can be lightened by drilling out part of the weight on the inboard end. The weight can be increased by a reasonable amount by ordering additional
weight and gang channel listed in the applicable Parts Catalog, and installing next to the
inboard weight the minimum amount necessary for correct balance. The minimum amount
that must be installed, however, must contain at least two attaching rivets. If this minimum
amount results in an over-balanced condition, the new weight and/or old weights can be
lightened.

Figure 18-3.
18-8

Control Surface Balancing (Sheet 1 of 5)

RUDDER AND ELEVATORS.
a. With the rudder/elevator set upon a FLAT, LEVELED surface, block up the trailing edge
until a center line through the attaching bolt and the trailing edge is equal distance from the
leveling surface (detail H).
b. Place the balancing beam (detail A) on the rudder/elevator near the center attaching bracket,
(90 ° to the hinge line). Adjust the trailing edge support on the balancing beam (detail H) until
the beam is level. If the rudder/elevator has not been disturbed during this operation, the
beam is now parallel to the chord line of the rudder/elevator.
NOTE
The above procedure must be performed with care. Small angular
discrepancies will produce large balancing errors.
c. Mark position of the balancing beam, then remove and balance the beam by itself at the knife
edges by adding washers as shown in (detail C).
d. Place the balancing beam on the rudder/elevator in its original position, then remove the block
from beneath the trailing edge.
e. Place the sliding weight (detail H) on the forward end of the balance beam, move it along the
beam until the beam is again level. A small, lightweight, spirit-level may be used for this
purpose provided it is symmetrical about its bubble reference and this reference is placed
on the beam directly over the rudder/elevator hinge line (detail H).
f. If the rudder/elevator is correctly balanced, the position of the sliding weight with respect to
the rudder/elevator hinge line, will produce a moment about the hinge line somewhere within
the underbalance tolerance listed in the chart on (Sheet 5 of 5).
g. If modification of the rudder/elevator balance weight is necessary to correct an out-of-balance
condition, the balance weight can be lightened by drilling out part of the weight. The weight
can be increased by fusing bar stock solder to the weight after removal from rudder/elevator.

BALANCING BEAM
Mark graduations in inches.

Four-foot length of extruded channelGrind weight to slide along beam, grind
ends to obtain exactly one pound, and
mark center of weight.Fabricate vertically adjustable
trailing edge support that will
slide along beam.

Attach knife edges and
mark at mid-point.

Detail

Figure 18-3.

Control Surface Balancing (Sheet 2 of 5)
18-9

1/16" SLOT: 3/4" DEEP

by adding washers and/

BALANCING

eail

SPIRIT-LEVEL

SUPPORT
BALANCING
MANDREL

AT AILERON
57

LEVELED SURFACE

10'

(W.S. 138. 00)

\AILERON

DetailD

*ALTERNATE METHOD
Befor making trailing edgel
measurement make sure

"d" - 1. 80
INCHES

aileron is straight in this
area.
AILERON-

PIANO HINGE

BALANCING MANDREL

Detail E

Figure 18-3.
18-10

Control Surface Balancing (Sheet 3 of 5)

A balance in this range is "underbalance".-

A balance in this range is "overbalance".

BALANCING MANDREL

RUDDER

Detail F

SPIRIT-LEVEL
PROTRACTOR

BALANCING
MANDREL

CHORD LINE

Detail H

Figure 18-3.

Control Surface Balancing (Sheet 4 of 5)
18-11

CONTROL SURFACE BALANCE REQUIREMENTS
NOTE
Unpainted values are not limits which must be met. They are given as guides, in order that the unbalance of the control surface in the final aircraft configuration may be predicted. If the control surface in the unpainted condition falls within the unpainted limit, the mechanic may feel confident that
the control surface will be acceptable after painting. However, if the surface in the unpainted condition exceeds the unpainted limit, the unbalance must be checked again after final painting to assure
that the control surface falls within the painted unbalance limit. Refer to GENERAL NOTES on sheet
3 for specific conditions.
DEFINITIONS:
UNDERBALANCE is defined as the condition that exists when the control surface is trailing edge heavy,
and is symbolized by a plus (+).
OVERBALANCE is defined as the condition that exists when the control surface is leading edge heavy,
and is symbolized by a minus (-).
NOTE
The "Balance Limits" columns list the moment tolerances within which the control surface
must balance. The tolerances must never be exceeded in the final flight configuration.
CONTROL: AILERON
UNPAINTED (Inch-Pounds)

PAINTED (Inch-Pounds)
BALANCE LIMITS

BALANCE LIMITS

0.0 to +9. 64

0.0 to+7.07

CONTROL:
PAINTED (Inch-Pounds)

RUDDER
UNPAINTED (Inch-Pounds)

BALANCE LIMITS

0.0 to +6. 0

0.0 to +4. 0

CONTROL: RIGHT ELEVATOR
UNPAINTED (Inch-Pounds)

PAINTED (Inch-Pounds)
BALANCE LIMITS

BALANCE LIMITS

0.0 to +20.47

0.0 to+18.1

CONTROL:
PAINTED (Inch-Pounds)
BALANCE LIMITS
0.0 to +20. 47

Figure 18-3.
18-12

Change 3

LEFT ELEVATOR
UNPAINTED (Inch-Pounds)
BALANCE LIMITS
0.0 to +18.1

Control Surface Balancing (Sheet 5 of 5)

PATCHES AND DOUBLERS 2024-T3 ALC LAD
4 REQD

DOUBLER

SECTION THRU PATCH
3. 00 DIA. HOLE

PATCH REPAIR FOR 3 INCH DIAMETER HOLE

MS20470AD4 RIVETS
16 REQD

22 1/20
4.00 DIA.

SKIN
2.00 DIA. HOLE

SECTION THRU PATCH
PATCH REPAIR FOR 2 INCH DIAMETER HOLE
2. 50 DIA.

SKIN
PATCH

DIA.

1.00 DIA. HOLE

SECTION THRU PATCH

ORIGINAL PARTS
REPAIR PARTS

OVERLAPPING

REPAIR PARTS IN CROSS SECTION

CIRCULAR PATCH

Figure 18-4.

Skin Repair (Sheet 1 of 6)
18-13

1/4 B

SECTION THRU ASSEMBLED PATCH

A-A
EDGE MARGIN = 2 X RIVET DIA.
PATCH - 2024-T3 ALCLAD

DIAMETER

RIVET SPACING =
6 X RIVET DIA.

RIVET TABLE
TANGULAR PATCH
.025
.032
.040
.051

ORIGINAL PARTS

.REPAIR PARTS

REPAIR PARTS IN CROSS SECTION

Figure 18-4.
18-14

Skin Repair (Sheet 2 of 6)

1/8
1/8
1/8
5/32

1/4 B

PATCH

EXISTING SKIN

pled skin and patch, and countersunk doubler.

SECTION THRU ASSEMBLED PATCH

s/--i/,

°,

NOTE

6 X RIVET DIA.

EDGE MARGIN =

2 X RIVET DIA.

DOUBLER -

2024-T3

. 032

1/8

REPAIR PARTS IN CROSS SECTION

Figure 18-4.

Skin Repair (Sheet 3 of 6)
18-15

NOTE
DOUBLER

DOUBLER
DOUBLER

Countersink doublers, and
dimple skin and patch.

EXISTING

DOUBLER

FLUSH PATCH AT
STRINGER/BU LKHEAD
INTERSECTION

NOTE

Figure 18-4.
18-16

Skin Repair (Sheet 4 of 6)

DOUBLERS-EXISTING

A-A
CARRY EXISTING
RIVET PATTERN

SECTION THRU ASSEMBLED PATCH

THRU PATCH

PITCH TYPICAL FOR

EDGE DISTANCE

2D MIN.

SPACER

PATCH

ORIGINAL PARTS
.

REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-4.

Skin Repair (Sheet 5 of 6)
18-17

-FUSELAGE
--SKIN

CLEAN OUT DAMAGED AREA

A-A
/°

PICK UP EXISTING
SKIN RIVET PATTERN

l0 RIVETS EACH SIDE OF

DAMAGED AREA
FILLER -2024-T4

ALCLAD

1/4" EDGE MARGIN-

ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-4.
18-18

Skin Repair (Sheet 6 of 6)

DOUBLER ALCLAD

2024-T4

-- DOUBLER 1/4"EDGE MARGIN

2024-T4 ALCLAD

STRINGER

6 RIVETS EACH SIDE --OF DAMAGED AREA

CLEAN OUT DAMAGED AREA

FILLER

2024-T4 ALCLAD

A-A

MS20470AD4 RIVETS

A
SKIN

ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-5.

Stringer and Channel Repair (Sheet 1 of 4)
18-19

FILLER -

2024-T4 ALCLAD

A-A

CLEAN OUT
DAMAGED AREA

RIVETS EACH SIDE

OF DAMAGED AREA

ANGLE -

2024-T4 ALCLAD
SPACING

STRINGER
PICK UP EXISTING SKIN RIVETS

MS20470AD4 RIVETS

ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-5.
18-20

Stringer and Channel Repair (Sheet 2 of 4)

STOP DRILL CRACK
A-A

1/4" EDGE MARGIN
SKIN

Figure 18-5.

Stringer and Channel Repair (Sheet 3 of 4)
18-21

FILLER

-2024-T4

A-A
CLEAN OUT DAMAGED AREA

SPACING

ORIGINAL PARTS
REPAIR PARTS
REPAIR IN CROSS SECTION

Figure 18-5.
18-22

Stringer and Channel Repair (Sheet 4 of 4)

ALCLAD

STOPDRILL CRACK IF CRACK
DOES NOT EXTEND TO EDGE
OF PART

DOUBLER2024-T3
ALCL.AD

ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-6.

Rib Repair (Sheet 1 of 2)
18-23

DOUBLER - 2024-T3 ALCLAD

-3/4"

RIVET

CLEAN OUT DAMAGED AREA

ANGLE - 2024-T4 ALCLAD

AROUND DAMAGED

ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN
CROSS SECTION
Figure 18-6.
18-24

Rib Repair (Sheet 2 of 2)

FILLER -

2024-T4 ALCLAD

DOUBLER--

CLEAN OUT DAMAGED AREA

2024-T4 ALCLAD

~:&

REPAIR PARTS

REPAIR PARTS N CROSS SECTION

Figure 18-7.

Wing Spar Repair (Sheet 1 of 4)
18-25

NOTE

N O.--

FILLER -

2024-T4 ALCLAD

This repair applies to either
front or rear spar if the spar
is a single channel.

CLEAN OUT DAMAGED AREA

7/8 x 7/8ALCLA
x.064

R2024-T4

1/4" EDGE MARGIN (TYP.)

ORIGINAL PARTS

REPAIR PARTS IN CROSS SECTION

Figure 18-7. Wing Spar Repair (Sheet 2 of 4)
18-26

3 ROWS RIVETS

A-A

FILLER ALCLAD

2024-T4

3/4" RIVET
SPACING

- CLEAN OUT

DAMAGED AREA
1/4" EDGE MARGIN--

A-A

MS20470AD4 RIVETS

ORIGINAL PARTS

REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-7.

Wing Spar Repair (Sheet 3 of 4)
18-27

FILLER
2024-0 ALCLAD
HEAT TREAT TO 2024-T4

2024-0 ALCLAD FILLERHEAT TREAT TO 2024-T4

CLEAN OUT

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

DAMAGED AREA

ORIGINAL PARTS

.........

REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-7.
18-28

18-28

STRIP-

Wing Spar Repair (Sheet 4 of 4)

A-A

DOUBLER -

Figure 18-8.

Auxiliary Spar Repair

2024-T4 ALCLAD

18-29

NOTES:
1.

Dimple leading edge skin and filler material; countersink the doubler.

Use MS20426AD4 rivets to install doubler.

Use MS20426AD4 rivets to install filler, except where bucking is impossible.
Cherry (blind) rivets where regular rivets cannot be bucked.

Contour must be maintained; after repair has been completed, use epoxy filler as necessary
and sand smooth before painting.

Vertical size is limited by ability to install doubler clear of front spar.

Lateral size is limited to seven inches across trimmed out area.

Number of repairs is limited to one in each bay.

Use CR162-4

1" MAXIMUM RIVET
SPACING (TYPICAL)

DOUBLER NEED NOT
BE CUT OUT IF ALL
RIVETS ARE ACCESSIBLE
FOR BUCKING

- .

5/16" MINIMUM EDGE
MARGIN (TYPICAL)

-- TRIM OUT DAMAGED AREA

FILLER MATERIAL

AS SKIN

Figure 18-9.
18-30

Leading Edge Repair

LEADING EDGE SKIN

1" MAXIMUM RIVET SPACING

1/4" MINIMUM EDGE MARGIN

TRIM OUT DAMAGED AREA

-FLAP LEADING EDGE SKIN
REPAIR DOUBLER TO BUTT
AGAINST CORRUGATED SKIN
TOP AND BOTTOM OF FLAP

.AT

1/4" MINIMUM EDGE MARGIN-

ALCLAD.

020

1/8" DIA. RIVETS

ORIGINAL PARTS
REPAIR PARTS

Figure 18-10.

Flap Leading Edge Repair
18-31

NOTES
Use rivet pattern at wing station

tion 23. 62 to wing station 85. 87.
Use rivet pattern at wing station

pattern at wing station 100. 00
with the number of BB4 dimpled
rivets at leading edge ribs between lap splices as shown:
*NO. OF CR2248
A.
*^XNO. OF CR2248-4
STATION *NO. OF BB4 RIVETS DIMPLED RIVETS
118
18
22
136
15
18
154
11
13
172
10
12
190
10
12

NO. OF CR2249-4
RIVETS

27
23
17
15
15
EXISTING
TACK RIVET

PATCH
EXISTING RIVET PATTERN
TYPICAL LEADING EDGE SECTION

NOTE
The Bulbed Cherrylock rivets listed may be substituted for BB4 dimpled rivets
in inaccessible areas, provided the number of rivets installed is increased proportionately. Blind rivets should not be installed in the wing spar.

Figure 18-11.
18-32

Change 2

Bonded Leading Edge Repair

USE EXISTING RIVET PATTERN

EXISTING AILERON SKIN

ORIGINAL PART
REPAIR PATCH IN CROSS SECTION

A-A
Figure 18-12.

Corrugated Skin Repair
18-33

LOWER WING SKIN (REF)

S-225-4F COVER

NOTE

MS20426AD3 RIVETS

PARTS ARE AVAILABLE
FROM THE CESSNA
S-1022Z-8-6

SCREWS

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

Add the minimum number of access holes necessary.
Any circular or rectangular access hole which is used with approved optional equipment installations may be added in lieu of the access hole illustrated.
Use landing light installations instead of adding access holes where possible. Do not add access
holes at outboard end of wing; remove wing tip instead.
Do not add an access hole in the same bay where one is already located.
Locate new access holes near the center of a bay (spanwise).
Locate new access holes forward of the front spars as close to the front spar as practicable.
Locate new access holes aft of the front spar between the first and second stringers aft of the
spar. When installing the doubler, rotate it so the two straight edges are closest to the stringers.
Alternate bays, with new access holes staggered forward and aft of the front spar, are preferable.
A maximum of five new access holes in each wing is permissible; if more are required, contact
the Cessna Service Department.
When a complete leading edge skin is being replaced, the wing should be supported in such a
manner so that wing alignment is maintained.
a.

Establish exact location for inspection cover and inscribe centerlines.

Determine position of doubler on wing skin and center over centerlines.
hole locations and drill to size shown.

Cutout access hole, using dimension shown.

Flex doubler and insert through access hole, and rivet in place.

Position cover and secure, using screws as shown.

Figure 18-13.
18-34

Access Hole Installation

Mark the ten rivet

A-A

CLEAN OUT DAMAGED AREA

ANGLE -

2024-T4 ALCLAD

10 RIVETS EACH SIDE

FIREWALL ANGLE
FILLER -

2024-T4 ALCLAD

FIREWA LL

1 ORIGINAL PARTS
REPAIR PARTS
REPAIR PARTS IN CROSS SECTION

Figure 18-14.

Firewall Angle Repair
18-35/(18-36 blank)

SECTION 19
PAINTING
NOTE
This section contains standard factory materials
listing and area of application. For paint number
and color, refer to Aircraft Trim Plate and Parts
Catalog. In all cases determine the type of paint
on the aircraft as some types of paint are not compatible. Materials may be obtained from the Cessna
Service Parts Center.
NOTE
The information in the following chart DOES NOT
apply to the A182 Series Aircraft.

CAUTION
When stripping aircraft of paint, use caution to
avoid stripper coming in contact with ABS parts.
MATERIAL

PAINT

NO /TYPE

ACRYLIC
LACQUER
LACQUER
EPOXY

PAINT

AREA OF APPLICATION

Used on exterior airframe.

Used on nose gear fairing on 1969 Models.

PRIMER

ER-7 WITH
ER-4
ACTIVATOR

Used with acrylic lacquer.

PRIMER

P60G2 WITH
R7K46
ACTIVATOR

Used with acrylic lacquer.

THINNER

T-8402A

THINNER

T-3871

Used with epoxy (Du Pont).

THINNER

T-6487

Used with epoxy (Enmar).

SOLVENT

#2 SOLVENT

Used to thin acrylic lacquer and for burndown.

Used to clean aircraft exterior prior to priming.

NOTE
Do not paint Pitot Tube, Gas Caps or Antenna covers
which were not painted at the factory.

Change 3

19-1

19-1. INTERIOR PARTS (Finish Coat of Lacquer)
a. Painting of Spare Parts.
1. Insure a clean surface by wiping with Naphtha
to remove surface contamination.
CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
2. After the part is thoroughly dry it is ready
for the lacquer topcoat. Paint must be thinned with
lacquer thinner and applied as a wet coat to insure
adhesion.
b. Touch Up of Previously Painted Parts.
1. Light sanding is acceptable to remove
scratches and repair the surface but care must be
exercised to maintain the surface texture or grain.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.
CAUTIONi
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
NOTE
Lacquer paints can be successfully spotted in.
19-2. EXTERIOR PARTS (Acrylic Topcoat)
a. Painting of Spare Parts.
1. Lightly scuff sand to remove scratches and
improve adhesion.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.

CAUTION
Do not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
3. After the part is thoroughly dry it is ready
for the topcoat. Paint must be thinned with appropriate acrylic thinner and applied as a wet coat to insure

19-2

Change 3

adhesion.
b. Touch Up of Previously Painted Parts.
1. Lightly scuff sand to remove scratches and
improve adhesion.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.
CAUTIONDo not use strong solvents such as Xylol,
Toluol or Lacquer Thinner since prolonged
exposure can soften or embrittle ABS.
3. Apply a compatible primer - surfacer and
sealer.
4. After the part is thoroughly dry it is ready
for the topcoat. Paint must be thinned and applied
as a wet coat to insure adhesion.
NOTE
Acrylic topcoats can be successfully spotted in.
19-3. EXTERIOR PARTS (Epoxy or Polyurethane
Topcoat)
a. Painting of Spare Parts and Touch Up of Painted
Parts.
1. Lightly scuff sand to remove scratches and
improve adhesion.
2. Insure a clean surface by wiping with Naphtha
to remove surface contamination.

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.

SECTION 20

WIRING DIAGRAMS

TABLE OF CONTENTS

Page

D. C. POWER
Battery and External Power Systems .
20-2
Battery and External Power Systems .
20-2A
Split Bus Bar. ............
20-3
60-Ampere Alternator ........
20-4
60-Ampere Alternator ........
20-5
60-Ampere Alternator ........
20-6
Split Bus Bar ............
20-7
IGNITION
Ignition System ...........
20-8
ENGINE CONTROL
Starter System ................
20-9
FUEL AND OIL
Oil Dilution .............
20-10
Oil Dilution .............
20-11
ENGINE INSTRUMENTS
Cylinder Head Temperature
Indicator ............
. 20-12
Fuel Quantity Indicator ........
20-13
Carburetor Air Temperature
Indicator ............
. 20-14
Hourmeter
.............
20-15
FLIGHT INSTRUMENTS
Turn and Bank and Gyro Horizon
Indicator .............
20-16
Brittain Wing Leveler. ........
20-17
Turn Coordinator
..........
20-18
Turn and Bank Indicator .......
20-19
Encoding Altimeter ..........
20-20
MISCELLANEOUS INSTRUMENTS
Clock ................
20-20A
LIGHTING
Dome and Courtesy Lights ......
. 20-21
Dome and Courtesy Lights .......
20-22

Instrument Lights and Compass ....
Landing Lights ............
Navigation Lights ..........
Flashing Beacon Light ........
Flashing Beacon Light ........
Map Light ..............
Electroluminescent Panel .......
Electroluminescent Panel .......
Electroluminescent Panel .......
Instrument and Oxygen ........
Instrument and Oxygen ........
Post Lighting ............
Post Lighting ............
Post Lighting ............
Post Lighting ............
Post Lighting ............
Control Wheel Map Light .......
Control Wheel Map Light ......
Control Wheel Map Light ......
Control Wheel Map Light .....
Control Wheel Map Light .....
Strobe Lights .
........
Strobe Lights ..
.......
Landing Lights
............
HEATING, VENTILATION AND DE-ICING
Cigar Lighter ............
Heated Pitot and Stall Warning
....
CONTROL SURFACES
Wing Flaps ............
Wing Flaps .............
Wing Flaps
.
........
Wing Flaps
.
........
WARNING AND EMERGENCY
Stall Warning (Non-heated). ......
Stall Warning (Non-heated .....

20-22A
20-23
20-24
20-25
20-26
20-27
20-28
20-29
20-30
20-30A
20-31
20-32
20-33
20-34
20-34A
20-35
20-36
. 20-37
. 20-38
. 20-38A
. 20-38B
.
20-39
.
20-40
20-40A
20-41
20-42
20-43
20-44
20-45
20-46

.
.

Change 3

20-47
20-48

20-1

CODE NO.

-XX

BTDO.N_

X|O.

SCALC:
I O4I

I PAoCI4.T

ON SHIELD.

SWITCH

NOTE:

rM3

S -15 799 -1
-I

BA.*T.CONITAC TOH
Q T5 ________

(OPT)

WHT

WLK

REVISIONS

REVISION

CONTRACT NO;

OI^~~~~~~~~

FORMNO

.0-211

S-XXX OR CMXXXX.CESSNA

SUPERSEDED BY:

OTHER

COMMERCIAL AIRCRAFT DIV.

DETAIL )

EE\f

W-

YEL(REF)

Y ZEV: AD S-Z160-l

,BEL

bLp
A

OH F-LA\k3

-& G-LA9

WIRE TABLE

NOTES :

l-_________________________BLK(I)
2.
-ZZ-0
O

So('t-?

CONTRACT NO:

PAWNEE DIVISION

INISTALL

S-IC.3'-r

TERNIlNAL

VerJDR

5 S- 3.-

Ia.

S5-.

HFou.5\,.

15
4OOLZ-Ss TEPM\ak_\

ebON
01>R

=5-1370-1 5i-

WIRE

I-________

TABLE

REVISIONS

[£

\2%

WAS 5-1637-1; 5-1641-9 WAS

20
Lazo

SRI.

"*a~~~~~~~~~

5~~~~~~~~S-XXX
OR CMXXXXCESSNA

'3-

I-I-

cJIT C.¥

WIRE TABLE

©A~m.RC

CA, 7

CKT

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Linearized                      : No
Creator                         : 
Producer                        : Avantext, Inc.
Modify Date                     : 2007:12:06 14:54:38-05:00
Create Date                     : 2001:07:23 09:53:01-04:00
Title                           : D2006-4-13 - MODEL 182 AND SKYLANE SERIES (1969 THRU 1976)
Subject                         : MODELS 182 AND SKYLANE SERIES (1969 THRU 1976)
Author                          : 
AVTX LLIB                       : MM
AVTX LPROD                      : CS03
AVTX P Notes                    : 1
Page Count                      : 387
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Metadata Date                   : 2007:12:06 14:54:38-05:00
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D2006 4 13 182 AND SKYLANE SERIES (1969 THRU 1976) Cessna_182_Skylane_1969 1976_D2006 Cessna 1969 1976

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