ARMY MODEL UH-1H/V HELICOPTERS. OPERATOR’S MANUAL (1988)

TM 58-1520-210-10

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OPERATOR’S MANUAL

ARMY MODEL

UH-1 H/V HELICOPTER

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ARMY MODEL

UH- IH/V HELICOPTERS

TM 55-1520-210-10, 15 February 1988, is changed as follows:

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Operator’s Manual

ARMY MODEL

UH-1HN HELICOPTERS

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By Order of the Secretary of the Army:

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Operator's Manual

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Operator’s Manual

ARMY MODEL

UH-lH/V HELICOPTERS

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Operator’s Manual

ARMY MODEL

UH-1H/V HELICOPTERS

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By Order of the Secretary of the Army:

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Official:

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MILTON H. HAMILTON

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23 July 1993

Operator's Manual

ARMY MODEL

UH-1H/V HELICOPTERS

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Operator's Manual

ARMY MODEL

UH-1H/V HELICOPTERS

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17 December 1990

Operator's Manual

Army Model

UH-1H/V Helicopters

TM 55-1520-210-10, 15 February 1988, is changed as follows:

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By Order of the Secretary of the Army:

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The Adjutant General

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WASHINGTON, D.C., 2 April 1990

Operator’s Manual

ARMY MODEL

UH-1H/V HELICOPTERS

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Operator's Manual

ARMY MODEL

UH-1H/V HELICOPTERS

TM 55-1520-210-10, 15 February 1988, is changed as follows:

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ARMY MODEL

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DATES OF ISSUE FOR ORIGINAL AND CHANGED PAGES ARE:

Original

15 February 1988

Change

18 April 1991

Change

28 July 1988

Change

01 July 1993

Change

17 October 1988

Change

23 July 1993

Change

09 December 1988

Change

30 November 1993

Change

15 March 1989

Change

09 March 1994

Change

11 December 1989

Change

14 February 1996

Change

02 April 1990

Change

13 August 1996

Change

17 December 1990

Change

13 February 1997

Change

27 December 1990

Change

08 October 1999

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2-36

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3-1

3-2 - 3-11

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1-1

3-13 - 3-21

1-2

3-22 - 3-22.28

2-1

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3-28 - 3-30

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3-31

2-5 - 2-8.1

3-32 - 3-32.24

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2-9 -2-12.1

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2-12.2 blank

3-39 - 3-40

2-13 - 2-16.1

3-41 - 3-42

2-16.2 blank

3-43 - 3-87

2-17-2-22

3-68 - 3-68.3

2-23 - 2-24

3-68.4 blank

2-25 - 2-26

3-69 - 3-90

2-26.1

4-1 - 4-4

2-26.2 blank

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2-27

4-6 blank

2-28

4-7 blank

2-29

4-8-4-10

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TM 55-1520-210-10

TECHNICAL MANUAL

HEADQUARTERS

DEPARTMENT OF THE ARMY

WASHINGTON D.C., 15 February 1988

OPERATOR'S MANUAL

ARMY MODEL

UH-1 H/V HELICOPTER

REPORTING ERRORS AND RECOMMENDING IMPROVEMENTS

You can help improve this publication. if you find any mistakes, or if you know of a way to improve the procedures,

please let us know. Mail your letter, DA Form 2028 (Recommended Changes to Publications and Blank Forms),

or DA Form 2028-2 located in the back of this manual direct to: Commander, U. S. Army Aviation and Missile Com-

mand, ATTN: AMSAM-MMC-LS-LP, Redstone Arsenal, AL 35898-5239. A reply will be furnished to you. You may

also send in your comments electronically to our e-mail address:

Is-lp@redstone.army.mil or fax

256-642-6546/DSN 788-6546. Instructions for sending an electronic 2028 may be found at the back of this manual

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Distribution Statement A. Approved for public release; distribution is unlimited.

TABLE OF CONTENTS

Page

CHAPTER 1

INTRODUCTION

1-1

CHAPTER 2

HELICOPTER AND SYSTEMS DESCRIPTION AND OPERATION

2-1

Section I

Helicopter

2-1

Emergency Equipment

2-3

ill

Engine and Related Systems

2-3

Helicopter Fuel System

2-6

Flight Control System

2-7

Hydraulic System

2-6

VII

Power Train System

2-9

VllI

Rotors

2-10

Utility Systems

2-10

Heating and Ventilation

2-11

Electrical Power Supply and Distribution System

2-12

XII

Lighting

2-13

xiii

Flight instruments

2-17

xiv

Servicing Parking and Mooring

2-19

CHAPTER 3

AVIONICS

3-1

Section I

General

3-1

Communication

3-1

Ill

Navigation

3-22

Transponder and Radar

3-32.24

CHAPTER 4

MISSION EQUIPMENT

4-1

Section I

Armament

4-1

Mission Avionics

4-5

Ill

Cargo Handling

4-5

Change 18

TM 55-1520-210-10

TABLE OF CONTENTS (CONT)

Page

CHAPTER 5

OPERATING LIMITS AND RESTRICTIONS

5-1

Section I

General

5-1

System Limits

5-1

III

Power Limits

5-1

Loading Limits

5-2

Airspeed Limits

5-2

Maneuvering Limits

5-2

VII

Environmental Restrictions

5-3

VIII

Height Velocity

5-3

Internal Rescue Hoist (Breeze Only)

5-3

Other Limitations

CHAPTER 6

WEIGHT/BALANCE AND LOADING

Section I

General

6-1

Weight and Balance

6-1

III

Fuel/Oil

6-2

Personnel

6-2

Mission Equipment

6-2

Cargo Loading

6-3

VII

Center of Gravity Limits

6-4

CHAPTER 7

PERFORMANCE DATA

Section I

Introduction

7-1

Torque Available

7-3

III

Hover

7-4

Takeoff

7-4

Cruise

7-5

Drag

7-6

VII

Climb-Descent

7-6

VIII

Fuel Flow

7-7

CHAPTER 7.1

PERFORMANCE DATA

Section I

introduction

7.1-1

Torque Available

7.1-3

III

Hover

7.1-4

Takeoff

7.1-4

Cruise

7.1-5

Drag

7.1-6

VII

Climb-Descent

7.1-6

VIII

Fuel Flow

7.1-7

CHAPTER 8

NORMAL PROCEDURES

Section I

Mission Planning

8-1

Crew Duties

8-1

III

Operating Procedures and Maneuvers

8-1

Instrument Flight

8-9

Flight Characteristics

8-9

8-11

Adverse Environmental Conditions

CHAPTER 9

EMERGENCY PROCEDURES

Section I

Helicopter Systems

9-1

Change 8

TM 55-1520-210-10

LIST OF ILLUSTRATIONS

FIGURE

TITLE

PAGE

2-1

General Arrangement Diagram

2-23

2-2

Principal Dimensions Diagram - Typical

2-25

2-3

Pilot/Copilot Seats - Typical

2-26

2-4

instrument Panel - Typical

2-26.1

2-5

Crew Compartment - Typical

2-28

2-6

Engine/Miscellaneous Control Panel Typical

2-31

2-7

Heating and Defrosting System

2-32

2-8

Electrical System Typical

2-34

2-9

Caution Panel - Typical

2-35

2-10

Servicing Diagram - Typical

2-36

3-1

Signal Distribution Panel SB-329/AR

3-2

Signal Distribution Panel C-1611/A/C

3-3

Signal Distribution Panel C6533/ARC

3-4

UHF Control Panel C6287/ARC-51BX

3-6

3-5

UHF Control Panel C-4677/ARC-51X

3-7

3-6

UHF Control Panel C-1827/ARC-55B

3-8

3-7

Control Panel AN/ARC-115

3-10

3-6

VHF Control Panel C-7197/ARC-134

3-11

3-9

VHF Control Panel 614U-6/ARC-73

3-13

3-10

Control Panel AN/ARC-114 and AN/ARC-114A

3-14

3-11

FM Radio Set Control Panel AN/ARC-131

3-15

3-12

FM Radio Set Control Panel C-3835/ARC-54

3-16

3-13

FM Control Panel and Switch Assembly AN/ARC-44

3-17

3-14

Voice Security Equipment

3-18

3-14.1

HF Radio Set AN/ARC-220 Control Panel C12436/URC

3-22.2

3-14.2

Remote Control Unit (TSEC/KY-100)

3-22.17

3-14.3

Processor (TSEC/KY-100)

3-22.21

3-15

HF Radio Control Panel

3-23

3-16

Direction Finder Control Panel ARN-83

3-24

3-17

ADF Control Panel ARN-59

3-26

3-18

Navigation Control Panel ARN-82

3-27

3-19

VHF Navigation Receiver Control Panel ARN-30E

3-27

3-20

Course Deviation Indicator ID-43/ARN-30 and ID-1347/ARN-82

3-28

3-21

Gyromagnetic Compass Indicator (RMI)

3-30

3-22

Marker Beacon Controls

3-31

3-22.1

Receiver/Display Unit

3-32.2

3-22.2

Instrument Panel - AN/ASN-175 System Components

3-32.7

3-23

DME indicator ID-2192/ARN-124

3-33

3-24

Transponder Set AN/APX-72

3-34

3-25

Transponder Set (AN/APX-100) Control Panel

3-35

3-26

AAU-32/A Altitude Encode/Pneumatic Altimeter

3-39

3-27

AN/APN-209 Radar Altimeter

3-39

3-28

Proximity Warning Panel

3-40

3-29

Radar Warning System

3-43

3-29.1

Receive r-Transmitter Radio RT-1167/ARC-164(V)

3-44

3-29.2

Control-Display Panel Layout

3-52

3-29.3

HF Volume Control

3-55

3-29.4

Channel Selection

3-56

3-29.5

Channel Stewing

3-57

3-29.6

Numbered Keys

3-57

3-29.7

Frequency Select

3-57

3-29.8

Frequency Slewing (RX only)

3-58

3-29.9

Power Level Select

3-59

3-29.10

Modulation Mode Source

3-59

3-29.11

Modulation Mode

3-60

3-29.12

Store

3-60

Change 18

iii

TM 55-1520-210-10

LIST OF ILLUSTRATIONS (CONT)

FIGURE

TITLE

PAGE

3-29.13

Selective Receive Address

3-60

3-29.14

Selective Transmit Address

3-61

3-29.15

Seladr ON/OFF

3-61

3-29.16

Test Mode Display

3-62

3-29.17

Failed Test Display

3-62

3-29.18

Passed Test Display

3-63

3-29.19

Tune Mode Single Channel

3-63

3-29.20

Tune Mode - All Channels

3-63

3-29.21

Scan Mode

3-63

3-29.22

Executing Scan

3-64

3-29.23

HF Volume Control

3-65

3-29.24

LF/ADF Control Panel C-7932/ARN-89

3-66

3-29.24.1

ADF Control Panel C-12192/ARN-149(V)

3-68.1

3-29.25

Radio Receiving Set AN/ARN-123(V)

3-69

3-29.26

Doppler Navigation Set AN/ASN-128

3-72

3-29.27

Doppler Lamp Test Mode Display

3-75

3-29.28

Horizontal Situation Indicator Control (C-11740/A)

3-84

3-29.29

Horizontal Situation Indicator (ID-2103/A)

3-85

3-29.30

Voice Security System Equipment

3-87

3-30

AN/ASC-15A(V) 1,2,3 and 4 Equipment Configuration

3-88

4-1

Mine Dispenser Control Panel - Typical

4-13

4-2

Hoist Installation - Typical

4-14

4-3

Pendant Control - Rescue Hoist

4-15

4-4

Hoist Cable Cutter Switch . Pilot - Typical

4-15

4-5

Hoist Cable Cut Switch - Hoist Operator - Typical

4-15

4-6

Control Pendant Assembly

4-16

4-7

AN/ASC-11 5A(V) 1, 2, 3 and 4 Equipment Configuration

4-17

5-1

Instrument Markings

5-4

5-2

Airspeed Operating Limits Chart

5-7

5-2.1

Airspeed Operating Limits Chart

5-8

6-1

Helicopter Station Diagram

6-5

6-2

Fuel Loading

6-6

6-3

Personnel Loading

6-8

6-4

Hoist Loading Limitations (Lateral CG)

6-9

6-5

Hoist Loading Limitations (Longitudinal CG)

6-10

6-6

System Weight and Balance Data Sheet

6-11

6-7

Hoist Installation Positions

6-14

6-8

Cargo Compartment

6-15

6-9

Cargo Tiedown Fitting Data

6-16

6-10

Internal Cargo Weight and Moment

6-17

6-11

External Cargo Weight and Moment

6-18

6-12

Center of Gravity Limits

6-19

Change 18

TM 55-1520-210-10

LIST OF ILLUSTRATIONS (CONT)

FIGURE

TITLE

PAGE

7-1

Temperature Conversion Chart

7-12

7-2

Maximum Torque Available (30 Minute Operation) Chart

7-13

7-3

Hover (Power Required) Chart

7-14

7-4

Control Margin

7-16

7-5

Takeoff Chart

7-18

7-6

Cruise Chart

7-21

7-7

Drag Chart

7-48

7-8

Climb-Descent Chart

7-50

7-9

Idle Fuel Flow Chart

7-51

7-1.1

Temperature Conversion Chart

7.1-8

7-1.2

Maximum Torque Available (30 Minute Operation) Chart

7.1-9

7-1.3

HoverChart

7.1-11

7-1.4

Control Margin Chart

7.1-15

7-1.5

Takeoff Chart

7.1-17

7-1.6

Cruise Chart

7.1-20

7-1.7

Drag Chart

7.1-46

7-1.8

Climb Chart

7.1-50

7-1.9

Fuel Flow Chart

1-52

8-1

Danger Area

8-15

8-2

Extenor Check Diagram

8-16

9-1

Emergency Exits and Equipment

9-10

9-2

Autorotational Glide Charactenstr-s Chart

9-11

9-2.1

Autorotational Glide Characteristics Chart

9-12

9-3

Height Velocity Diagram

9-13

9-3.1

Height Velocity Diagram

9-14

Change 16

v/(vi blank)

TM 55-1520-210-10

Operator’s Manual for

UH-1H/V Helicopters

Placement of tables, figures, and appendixes. Full

Distribution. Special distribution of this issue has been

page tables, figures, and appendixes (in that order) in-

made in accordance with DA Form 12-31-R Require-

cluded in this UPDATE printing are located following the

ments For Technical Manuals. The number of copies re-

chapters in which they were referenced.

quested in Block 275 and 668 of the subscribers DA

Form 12-31-R.

User Information. The proponent agency of this manu-

al is the US Army Aviation and Missile Command (AM-

COM). Submit changes for improving this publication on

Resupply. Limited copies of this publication are avail-

DA Form 2028 (Recommended Changes to Publica-

able from the St. Louis Publications Center for emer-

tions and Blank Forms) to: Commander, U.S. Army Avi-

gency requirements. Users should take action to insure

ation

and

Missile

Command

ATTN:

they are receiving the correct number by updating their

AMSAM-MMC-LS-LP, Redstone Arsenal, AL

DA 12-Series Forms.

35898-5230.

Types of comments that should be avoided on DA Form

2028 are those that: (1) Ask a question instead of giving

Explanation of Change Symbols. Changes in text

an answer; (2) Are based on minor differences of opin-

material and tables will be shown by a bar in the left mar-

ion or wording; (3) Point out obvious editorial errors,

gin. Correction of minor inaccuracies such as spelling,

mispellings, or errors in punctuation, unless the errors

punctuations, relocation of material, etc. will not be

change the intended meaning. (Reference AR 25-30,

shown unless such corrections changes the meaning of

Chapter 2).

instructive information and procedures.

Change

1-1

TM 55-1520-210-10

Chapter 1

Introduction

1-1. General These instructions are for use by the

1-5.

Appendix B, Abbreviations and Terms

operator(s). They apply to UH-1H/V helicopter.

Definitions of all abbreviations and terms used

throughout the manual are included in appendix B.

1-2. Warnings, Cautions, and Notes Warnings,

Cautions, and Notes are used to emphasize

1-6. Index The index lists every titled paragraph

important and critical instructions and are used for

contained in this manual. Chapter

7 performance

the following conditions.

data has an additional index within the chapter.

WARNING

1-7.

Army Aviation Safety Program_Reports

necessary to comply with the safety program are

An operating procedure, practice, etc.,

prescribed in AR

385-40.

which if not correctly followed, could re-

sult in personal injury or loss of life.

1-8.

Destruction of Army Materiel to Prevent

Enemy Use For information concerning destruction

of Army materiel to prevent enemy use, refer to TM

Caution

750-244-1-5.

An operating procedure, practice, etc.,

1-9.

Forms and Records

Army aviators flight

which if not strictly observed, could result in

records and helicopter maintenance records which

damage to or destruction of equipment.

are to be used by the operators and crew members

NOTE

are prescribed in DA PAM

738-751 and TM

55-1500-342-23

An operating procedure, condition, etc.,

1-10. Designator Symbols Designator Symbols H

which it is essential to highlight.

UH-1H and V UH-1V are used in conjunction with text

contents, text headings and illustration titles to show

1-3. Description This manual contains the best

limited effectivity of the material. One or more designator

operating instructions and procedures for the

symbols may follow a text heading or illustration title to in-

UH-1H/V helicopter under most circumstances. The

dicate proper effectivity, unless the material applies to all

observance of limitations, performance and weight

models and configurations within the manual. If the

balance

data provided is mandatory. The

material applies to all models and configurations, no

observance of procedure is mandatory except when

designator symbols will be used. The Designator Sym-

modification is required because of multiple

bols used for the Composite Main Rotor Blade is

emergencies, adverse weather, terrain, etc. Your

and the Metal Main Rotor Blade is

flying experience is recognized, and therefore, basic

flight principles are not included. THIS MANUAL

1-11. Use of Words Shall, Should, and May Within

this technical manual, the word “shall” is used to indicate

SHALL BE CARRIED IN THE HELlCOPTER AT ALL

a mandatory requirement. The word “should” is used to

TIMES.

indicate a non-mandatory but preferred method of ac-

complishment. The word “may” is used to indicate an ac-

1-4. Appendix A, References Appendix A is a

ceptable method of accomplishment.

listing of official publications cited within this manual

applicable to and available for flight crews.

1-2

Change 8

TM 55-1520-210=10

Chapter 2

Helicopter and Systems Description and Operation

Section 1.

HELlCOPTER

2-1. General Description The UH-1H/N helicopters

b. Tail Skid. A tubular steel tail skid is installed on

are thirteen-place single engine helicopters. The

the aft end of the tailboom. It acts as a warning to

maximum gross weight is 9500 pounds.

the pilot upon an inadvertent tail-low landing and

aids in protecting the tail rotor from damage.

2-2. General Arrangement Figure 2-1 depicts the

general arrangement. Indexed items include access

2-8.

Crew Compartment Diagram

The crew

openings and most of the items referred to in the

compartment is depicted in figure 2-5.

exterior check paragraph in section Ill of chapter

2-9. Cockpit and Cabin Doors

2-3. Principal Dimensions Figure 2-2 depicts the

a. Cockpit Doors. The cockpit doors are formed

principal dimensions.

aluminum frames with transparent plastic windows in

2-4. Turning Radius Turning radius is about

the upper section (fig 2-1). Ventilation is supplied by

feet when pivoted around the mast.

the sliding panels in the windows. Cam-type door

latches are used and doors are equipped with jet-

2-5. Fuselage The fuselage is the forward section

tisonable door releases.

of the airframe extending from the nose to the

forward end of the tailboom. The fuselage consists

b. Cabin Doors. The two cabin doors are formed

primarily of two longitudinal beams with transverse

aluminum frames with transparent plastic windows in

bulkheads and metal covering. The main beams are

the upper section (fig 2-1). These doors are on roll-

the supporting structure for the cabin, landing gear,

ers and slide aft to the open position allowing full

fuel tanks, transmission, engine, and tailboom. The

access to the cargo area. Hinged doorpost panels

external cargo suspension unit is attached to the

are forward of the cabin doors. They provide a

main beams near the center of gravity of the

larger entrance to the cargo area. An open door

helicopter.

lock is provided to hold the door in the aft position to

prevent door separation in flight.

2-6. Tailboom The tailboom section is bolted to

the aft end of the fuselage and extends to the aft

2-10.

Pilot/Copilot Seats The pilot and copilot

end of

the

helicopter.

a tapered,

seats may be conventional seats or armored seats

semi-monocoque structure comprised of skins,

(fig

2-3). The armored seats have a release to

Iongerons, and stringers. The tailboom supports the

recline the seats to aid in removal of injured

tail rotor, vertical fin, and synchronized elevator. It

personnel. The conventional seats do not have the

houses the tail rotor driveshaft and some electronic

reclining feature.

equipment.

a. Pilot and Copilot Seats (Conventional). The pilot

2-7. Landing Gear System

and copilot seats are vertical and fore-aft adjustable

and the nonreclining type. The vertical height adjust-

a. Main Landing Gear. The main landing gear con-

ment handle is under the right side of the seat. The

sists of two aluminum arched crosstubes mounted

fore and aft adjustment is under the left side of the

laterally on the fuselage with two longitudinal skid

seat. Webbing on the back of the seat can be re-

tubes attached to the crosstubes. The skid tubes

moved to accept use of a back-pack parachute.

are made of aluminum and have steel skid shoes

The seats are equipped with lap safety belts and in-

attached to the bottom to minimize skid wear.

ertia reel shoulder harness.

2-1

TM 55-1520-210-10

b. Plot and Copilot Seats (Armored). Armored

the approved patient securing straps when the helicopter

seats can be installed m the helicopter for the pilot and

is used for medical evacuation missions. Instruments

copilot. They are equipped with lap safety belt and

and Controls.

inertia-reel shoulder harness They are adjustable fore

and aft and vertically. The vertical adjustment handle Is

2-12 Instruments and Controls

under the right side of the seat and the fore and aft

handle on the left. The seats are equipped with a quick

a. Instrument Panel.

The location of all the

release, on each side at the back of the seat, for

controls, indicators, instruments, and data placards

reclining the seat. The seat back, bottom, and sides are

installed on the instrument panel is depicted m figure 24

protected by ceramic and aluminum armor plate. Hip

V Some instruments may be relocated.

and shoulder areas are protected by ceramic type armor.

b. Pedestal Panel.

The panels and controls

c. Inertia Reel Shoulder Harness. An inertia reel

installed in the pedestal are depicted in figure 2-5.

and shoulder harness is incorporated in the pilot and

copilot seats with manual lock-unlock handle (fig 2-3).

c. Overhead Console. The location of the controls

On the conventional seat, the control handles are located

and circuit breakers installed in the overhead console is

on the left front of the seat. On the armored seat, the

depicted in figure 2-5.

control handles are located on the night front of the seat.

With the control in the unlocked position (aft) and the

d. External Stores Jettison Handle. The external

shoulder straps properly adjusted, the reel strap will

stores Jettison handle Is located to the left of the pilot

extend to allow the occupant to lean forward; however,

collective when installed. Pulling up on the handle will

the reel automatically locks when the helicopter

Jettison external stores through mechanical linkage.

encounters an impact force of 2 to 3 ’G’ deceleration.

The reel can be locked (handle forward) from any

e. Other Instruments and Controls. Instruments,

position and will take up slack in the harness. To release

controls, or indicators not shown in figure 2-5 or figure 2-

the lock, it is necessary to lean back slightly to release

6 are shown in the chapter/section which describes their

tension on the lock and move the control handle to the

related systems.

unlock position. It is possible to have pressure against

the seat back whereby no additional movement is

2-12.1 Wire Strike Protection System (WSPS).

possible and the lock cannot be released. If this

condition occurs. It will be necessary to loosen the

The WSPS provides protection for 90 % of the frontal

harness. The reel should be manually locked for

area against impacts with horizontally strung mechanical

emergency landing. Straps must be adjusted to fully

and power transmission cables. The basic system

retract within the inertia reels to prevent rebound

consists of an upper cutter/deflector, a windshield

overshoot in the event of impact. Seat belt must be

protector/deflector/cutter, a lower cutter/deflector and a

securely fastened and firmly tightened prior to

pair of windshield wiper deflectors (fig 2-1). The lower

adjustment of shoulder harness to prevent submanning

cutter assembly features a ’Breakaway Tip’ designed to

in event of Impact.

shear when relatively large ground contact forces are

experienced and before helicopter structural damage is

2-11. Personnel Seats. Various arrangements of

incurred. However, the tip shear rivets are designed to

personnel seats can be installed to accommodate from

withstand the smaller forces experienced during wire

one to eleven personnel besides the pilot and copilot.

strikes and the tip will still effectively deflect wires/cables

The seats are constructed of tubular steel and reinforced

into the cutter blades.

canvas. Each seat is equipped with a lap safety belt.

For additional information on the personnel seats, refer

to chapter 6. Patients will be secured to litters utilizing

2-2

Change 17

TM 55-1520-210-10

Section II. EMERGENCY EQUIPMENT

2-13.

Emergency Equipment. The emergency

2-15. First Aid Kits. Four general purpose type first aid

equipment location, illustration, and emergency

kits have been provided in the cabin area (fig 9-1) Two

procedures are covered in chapter 9.

kits are secured to the night center doorpost. The other

two kits are secured to the left doorpost. First aid kits

2-14. Portable Fire Extinguisher. A portable hand-

can be easily removed for immediate use.

operated fire extinguisher is carried in a bracket located

aft of the pedestal, or to the right of the pilot seat.

Section III. ENGINE AND RELATED SYSTEMS

2-16. Engine. The UH-IH/ are equipped with a T53-L-

13 engine.

NOTE

The ice detector system is not applicable

2-17. Engine Compartment Cooling. The engine

on helicopters equipped with the self-

compartment is cooled by natural convection through

purging particle separator.

engine compartment screens.

d. DE-ICE. Engine de-ice is a bleed air system

2-18. Air Induction System. Three different air

activated by the DE-ICE switch on the ENGINE panel (fig

induction systems are used on these helicopters. They

2-6). In the ON position bleed air is directed through the

are discussed m the following paragraphs:

engine inlet to provide the protection. Power losses

caused when the system is on are shown m chapter 7.

a Non-Self Purging Particle Separator. The

In the event of dc electrical failure or when the DE-ICE

non-self-purging particle separator is an inertial type. A

ENG circuit breaker is out de-ice is automatically on.

lip extending into the airstream deflects the particle-laden

System power is provided by the dc essential bus and

air into a large chamber. Large particles in the air settle

protected by the ANTI-ICE ENG circuit breaker.

in the chamber; fine particles are removed as the air is

drawn through a filter assembly. Removed particles are

e. Improved Particle Separator Some UH.-l’s may

held m porous foam box assembles. The box

be equipped with an improved particle separator. This

assemblies can be removed and cleaned. Other

unit has a number of vortex tubes which ar-highly

components used with the particle separator are: ENG

effective in removing sand and dust from the engine inlet

AIR FILTER CONT circuit breaker on the overhead

air. The sand and dust are dumped overboard through

console, an engine air differential pressure switch on the

outlets on each side of the Separator.

firewall, and an cateye type on the instrument panel or of

the segment type on the caution panel.

2-19. Engine Fuel Control System.

b. Self-Purging Particle Separator.

Helicopters

a Engine Mounted Components. The fuel control

serial No.

68-15779 and subsequent are equipped with

assembly is mounted on the engine. It consists of a

a self-purging particle separator. This is an inertial-type

metering section a computer section and an overspeed

separator. Particle-laden air is directed through a large

governor.

annular chamber and through an air cleaner. A constant

supply of bleed air from the engine flows through the

(1) The metering section is driven at a speed

venture-type ejector and carries particles overboard

proportional to N1 speed. It pumps fuel to the engine

through airframe plumbing. Some self-purging particle

through the main metering valve or if the main system

separator systems have operational ENGINE INLET AIR

falls through the emergency metering valve which is

caution panel segment lights.

positioned directly by the twist grip throttle.

c Foreign Object Damage Screen Foreign Object

Damage (FOD) screen prevents large particles from

entering the engine inlet.

Change 17

2-3

TM 55-1520-210-10

(2) The computer section determines the rate

compressor stall, overspeed, over-temperature, or

of main fuel delivery by biasing main metering valve

engine failure. The governor circuit receives power from

opening for N I speed, inlet air temperature and

the 28 Vdc essential bus and is protected by the GOV

pressure, and throttle position. It also controls the

CONT circuit breaker.

operation of the compressor air bleed and operation of

the variable inlet guide vanes.

2-20. Engine Oil Supply System.

(3) The overspeed governor is driven at a

a. Description. The dry sump pressure type oil

speed proportional to N2 speed. It biases the main

system is entirely automatic in its operation. The system

metering valve opening to maintain a constant selected

consists of an engine oil tank with de-aeration provisions.

N2 rpm.

thermostatically controlled oil cooler with by-pass valve,

pressure transmitter and pressure indicator, low

b. Starting Fuel Flow. During engine start.

pressure warning switch and indicator, sight gages, and

energizing the start fuel switch opens the fuel solenoid

oil supply return vent, and breather lines. Drain valves

valve, allowing fuel from the fuel regulator to flow through

have been provided for draining the oil tank and cooler.

the starting fuel manifold and into the combustion

Pressure for engine lubrication and scavenging of return

chamber. When N1 reaches sufficient speed, the start

oil are provided by the engine-mounted and engine-

switch is de-energized, causing the solenoid valve to

driver oil pump. On helicopters equipped with Oil Debris

close and stop-starting fuel flow. Starting fuel nozzles

Detection System (ODDS), and external oil separator,

are purged by air from the combustion chamber through

with integral chip detector, and a 3-micron filter are

a check filter valve. Engine starting fuel solenoid valve is

installed down stream of the sump. Oil specification and

controlled by the engine starter switch on helicopters

grade are specified in the Servicing Table 2-1

which do not have a starting fuel switch. The engine

solenoid valve (engine starting fuel solenoid valve)

b. Oil Cooler. Engine oil cooling Is accomplished

cannot be individually controlled during engine starts.

by an oil cooler. The cooler is housed within the

fuselage area under the engine deck (fig 2-1). Air

c. Power Controls (Throttles). Rotating the pilot or

circulation for oil cooling is supplied by a turbine fan

copilot twist grip-type throttle (fig 2-5) to the full open

which operates from turbine bleed air. The fan is

position allows the overspeed governor to maintain a

powered at all times when the engine is operating and no

constant rpm. Rotating the throttle toward the closed

control is required except the bleed air limiting office.

position will cause the rpm to be manually selected

instead of automatically selected by the overspeed

2-21. Ignition Starter System. The starter ignition

governor. Rotating the throttle to the fully closed position

switch is mounted on the underside of the pilot collective

shuts off the fuel. An idle stop Is incorporated in the

pitch control lever switch box. An additional switch may

throttle to prevent inadvertent throttle closure. To bypass

be installed on the copilot stick. The switch is a trigger

the idle detent press the IDLE REL switch and close the

switch, spring-loaded to the off position (fig 2-5). The

throttle. The IDLE REL switch is a momentary on,

starter and ignition unit circuits are both connected to the

solenoid-operated switch. The IDLE REL switch is

trigger switches. The circuits receive power from the 28

located on the pilot collective stick switch box. IDLE REL

Vdc essential bus and are protected by circuit breakers

switch receives power from the 28 Vdc bus and is

marked STARTER RELAY and IGNITION SYSTEM

protected by a circuit breaker marked IDLE STOP REL.

IGNITER SOL. The starter circuit is energized when the

Friction can be induced in both throttles by rotating the

STARTERJGEN switch is in the START position and the

pilot throttle function ring counterclockwise (fig 2-5). The

trigger switch is pulled (fig 2-5).

ring is located on the upper end of the pilot throttle.

The ignition circuit is energized when the FUEL MAIN

d. Governor switch. The GOV switch is located on

ON/OFF switch on the engine control panel is m the ON

the ENGINE control panel (fig 2-6). AUTO position

position and the trigger switch is pulled. The ignition

permits the overspeed governor to automatically control

keylock is located by the AC circuit breaker panel. The

the engine rpm with the throttle in the full open position.

OFF position deactivates the igniters and start fuel to

The EMER position permits the pilot or copilot to

prevent engine starting. The ON position allows engine

manually control the rpm.

Because automatic

starting.

acceleration, deceleration, and overspeed control are not

provided with the GOV switch in the EMER position,

control movements must be smooth to prevent

2-4

Change 17

TM 55-1520-210-10

2-22. Governor RPM Switch. The pilot and copilot

The tachometer inner scale is marked ROTOR and the

GOV RPM INCR/DECR switches are mounted on a

outer scale is marked ENGINE. Synchronization of the

switch box attached to the end of the collective pitch

ENGINE and ROTOR needles indicates normal

control lever (fig 2-5). The switches are a three-position

operation of helicopter. The indicator receives power

momentary type and are held in INCR (up) position to

from the tachometer generators mounted on the engine

increase the power turbine (N2) speed or DECR (down)

and transmission. Connection to the helicopter electrical

position to decrease the power turbine (N2) speed.

system is not required.

Electrical power for the circuit is supplied from the 28

d. Gas Producer Tachometer. The gas producer

Vdc essential bus and is protected by a circuit breaker

indicator is located in the right center area of the

marked GOV CONT.

instrument panel and is marked PERCENT (fig 24). The

indicator displays the rpm of the gas producer turbine

2-23. Droop Compensator. A droop compensator

speed in percent. This system receives power from a

maintains engine rpm (N2) as power demand is

tachometer generator which is geared to the engine

increased by the pilot. The compensator is a direct

compressor. A connection to the helicopter electrical

mechanical linkage between the collective stick and the

system is not required.

speed selector lever on the N2 governor. No crew

e. Oil Temperature Indicator. The engine oil

controls are provided or required. The compensator will

temperature indicator is located in the center area of the

hold N2 rpm to j40 rpm when properly rigged. Droop is

instrument panel and is marked OIL ’C (fig 24). The

defined as the speed change in engine rpm (N2) as

indicator is connected to an electrical resistance-type

power is increased from a no-load condition. It is an

thermocouple. The temperature of the engine oil at the

inherent characteristic designed into the governor

engine oil inlet is indicated in degrees celsius. Power to

system. Without this characteristic instability would

operate the circuit is supplied from the 28 Vdc essential

develop as engine output is increased resulting in N I

bus. Circuit protection is provided by the TEMP IND

speed overshooting or hunting the value necessary to

ENG & XMSN circuit breaker.

satisfy the new power condition. If N2 power were

f. Oil Pressure Indicator The engine oil pressure

allowed droop other than momentarily the reduction in

indicator is located in the center area of the instrument

rotor speed could become critical.

panel and is marked OIL PRESS (fig 2-4). The indicator

receives pressure indications from the engine oil

2-24. Engine Instrument and Indicators. All engine

pressure transmitter and provides readings in pounds

instruments and indicators are mounted in the instrument

per square inch (psi). The circuit receives electrical

panel and the pedestal (figs 24 and 2-5).

power from the 28 VAC bus and circuit protection Is

provided by the ENG circuit breaker in the ac circuit

a. Torquemeter Indicator.

The torquemeter

breaker panel.

indicator is located in the center area of the instrument

g. Oil Pressure Caution Light The ENGINE OIL

panel and is marked TORQUE PRESS (fig 24). The

PRESS caution light Is located in the pedestal mounted

indicator is connected to a transmitter which is part of the

CAUTION panel. The light is connected to a low

engine oil system. The torquemeter indicates torque in

pressure switch.

When pressure drops below

pounds per square inch (psi) of torque imposed upon the

approximately 25 psi the switch closes an electrical

engine output shaft. The torquemeter receives power

circuit causing the caution light to illuminate. The circuit

from the 28 Vac bus and is protected by a circuit breaker

receives power from the 28 Vdc essential bus and is

marked TORQUE in the ac circuit breaker panel b.

protected by the circuit breaker marked CAUTION

Exhaust Gas Temperature Indicator. The exhaust gas

LIGHTS.

temperature indicator is located in the center area of the

h. Engine Chip Detector Caution Light. A magnetic

instrument panel and is marked EXH TEMP (fig 24).

plug is installed in the engine. When sufficient metal

The indicator receives temperature indications from the

particles accumulate on the magnetic plug to complete

thermocouple probes mounted in the engine exhaust

the circuit the ENGINE CHIP DET segment illuminates.

diffuser section. The temperature indications are in

The circuit receives power from the 28 Vdc essential bus

degrees celsius.

The system is electrically self-

and is protected by the circuit breaker marked CAUTION

generating.

LIGHTS. On Helicopters equipped with ODDS, the chip

c. Dual Tachometer. The dual tachometer is

detector which is connected to the caution light is part of

located in the center area of the instrument panel and

the external oil separator.

indicates both the engine and main rotor rpm (fig 2 4).

Change 17

2-5

TM 55-1520-210-10

i. Engine Ice Detector. The ice detector system

(ENGINE ICE DET caution light) is not connected.

m. Failure of either fuel pump element will close an

electrical circuit illuminating the caution light. The system

j. Engine Icing Caution Light The ENGINE ICING

receives power from the 28 Vdc essential bus and is

segment of the caution panel is not connected.

protected by a circuit breaker marked CAUTION

LIGHTS. One type of switch used on some aircraft will

illuminate the caution light until normal operating

NOTE

pressure is reached. This momentary lighting does not

Engine inlet air filter clogged/caution

indicate a pump element failure.

lights are not utilized for aircraft with the

improved particle separator. However,

n. Emergency Fuel Control Caution Light The

some self-purging particle separators

emergency fuel control caution light is located m the

have operational ENGINE INLET AIR

pedestal-mounted caution panel. The Illumination of the

segment lights.

worded segment GOV EMER Is a remainder to the pilot

that the GOV switch is in the EMER position. Electrical

k Engine Inlet Air Filter Clogged Warning Light On

power for the circuit is supplied from the 28 Vdc bus and

helicopters pnor to Serial No.

68-16066 the ENGINE

is protected by a circuit breaker marked CAUTION

INLET FILTER CLOGGED warning light is mounted on

LIGHTS.

the upper area of the instrument panel (fig 2-4). When

the inlet air filter becomes clogged, a differential

o. Fuel Filter Caution Light The FUEL FILTER

pressure switch senses the

condition and closes

caution light Is located in the pedestal-mounted caution

contacts to energize the filter caution light. Power is

panel or a press to test light is located on the instrument

supplied from the 28 Vdc bus and circuit protection is

panel. A differential pressure switch is mounted m the

provided by the CAUTION LIGHTS circuit breaker.

fuel line across the filter. When the filter becomes

clogged, the pressure switch senses this and closes

i Engine Inlet Air Caution Light The ENGINE INLET AIR

contacts to energize the caution light circuit. If clogging

segment of the caution panel will illuminate when the

continues, the fuel bypass opens to allow fuel to flow

inlet air filter becomes clogged. Power is supplied from

around the filter. The circuit receives power from the 28

the

28 Vdc bus and protection is provided by the

Vdc essential bus and Is protected by a circuit breaker

CAUTION LIGHT circuit breaker.

marked CAUTION LIGHTS.

Section IV. HELICOPTER FUEL SYSTEM

2-25. Fuel Supply System.

severe, but survivable, crash Impact to reduce the

possibility of fire. Frangible fittings used to secure the

a. Fuel System The fuel system consists of five

fuel cells m the airframe are designed to fuel and permit

interconnected cells all filled from a single fitting on the

relative movement of the cells, without rupture, in event

right side of the helicopter. The two forward cells each

of a crash; self-sealing break-away valves are installed in

contain a submerged boost pump. The boost pumps

the fuel lines at the fuel cell outlets and certain other

provide fuel pressure to prime the fuel line to the engine

locations. The break-away valves are designed to permit

driven fuel pump. Each forward fuel cell is divided into

complete separation of components without loss of fuel.

two compartments by a lateral baffle fitted with a flapper

Rollover vent valves are installed on the aft fuel cells to

valve to allow fuel flow from front to rear.

The

provide protection m the event of a helicopter rollover

submerged boost pump is mounted on a sump assembly

during a crash. The system has 50 caliber ballistic

near the aft end of each forward cell and is connected by

protection in the lower two-thirds of the cell.

a hose to the pressure line outlet to the engine. Part of

the pump output is diverted forward through a flow switch

b. Closed Circuit Refueling System Helicopter

and hose to an ejector pump at front of cell. Induced flow

serial number 69-15292 and subsequent and modified

of the ejector pump sends fuel through a hose over the

helicopters provide a closed circuit refueling system

baffle into the rear part of the cell, so that no slgnmficant

when used with the mating nozzle. This system is

quantity of fuel will be unusable in any flight attitude. The

capable of automatic shut-off of fuel flow when full.

crashworthy system is designed to contain fuel during a

2-6

Change 17

TM 55-1520-210-10

c Gravity Refueling.

If helicopter is equipped with

transistorized electrical receiver

which

closed circuit refueling system and fuel servicing vehicle

continuously indicates the quantity of fuel m pounds.

is not equipped with related nozzle for closed circuit

The indicator is connected to three fuel transmitters

refueling, a gravity system may be used.

mounted m the fuel cells. Two are mounted in the night

forward cell and one in the center aft cell. Indicator

2-26. Controls and Indicators.

readings shall be multiplied by 100 to obtain fuel quantity

m pounds. Electrical power for operation is supplied

a Fuel Switches.

The fuel system switches

from the 115 Vac system and is protected by circuit

consist of a main fuel switch, start fuel switch, and fuel

breaker FUEL QTY in the ac circuit breaker panel.

transfer switches (fig 2-6). The FUEL START switch is

not applicable on helicopters Serial No. 66-8574 through

c Fuel Gage Test Switch The FUEL GAGE TEST switch

66-8577, 66-16034 and subsequent, and earlier models

is used to test the fuel quantity indicator operation (fig 2-

so modified.

4).Pressing the switch will cause the indicator pointer to

move from the actual reading to a lesser reading.

(1) Main Fuel Switch The FUEL MAIN

Releasing the switch will cause the pointer to return to

ON/OFF switch is located on the pedestal-mounted

the actual reading. The circuit receives power from the

ENGINE panel (fig 2-6). The switch is protected from

115 Vac system and is protected by a circuit breaker

accidental operation by a spring-loaded toggle head that

marked FUEL QTY m the ac circuit breaker panel.

must be pulled up before switch movement can be

accomplished. When the switch is in the ON position,

the fuel valve opens, the electric boost pump(s) are

d Fuel Pressure Indicator The fuel pressure

energized and fuel flows to the engine. When the switch

indicator displays the psi pressure of the fuel being

is in the OFF position the fuel valve closes and the

delivered by the boost pumps from the fuel cells to the

electric boost pump(s) are de-energized. Electrical

engine (fig 2-4). The circuit receives power from the 28

power for circuit operation is supplied by the 28 Vdc

Vac bus and is protected by the FUEL PRESSURE

essential bus and is protected by circuit breakers FUEL

circuit breaker in the ac circuit breaker panel.

VALVES, LH BOOST PUMP and RH ROOST PUMP.

e. Fuel Quantity Low Caution Light The

(2) Fuel Start Switch. The FUEL START ON/OFF

MINUTE FUEL caution light will illuminate when there is

switch is located on the ENGINE panel. In the ON

approximately 170 pounds remaining. The illumination

position the starting fuel solenoid valve is energized

of this light does not mean a fixed time period remains

when the starter-ignition switch is pulled. When the

before fuel exhaustion, but is an medication that a low

START FUEL switch is in the OFF position the igniter

fuel condition exists. Electrical power is supplied from

solenoid valve is de-energized. Electrical power for the

the 28 Vdc essential bus. The caution lights circuit

circuit is supplied by the 28 Vdc essential bus and is

breaker protects the circuit.

protected by circuit breaker IGNMTION SYSTEM

IGNITER SOL.

NOTE

Low fuel caution systems alert the pilot

(3) Fuel Control.

Fuel flow and mode of operation

that the fuel level in the tank has

is controlled by switches on the pedestal-mounted

reached a specified level (capacity).

engine control panel (fig 26). The panel contains the

Differences in fuel densities due to

MAIN FUEL ON/OFF or FUEL ON/OFF switch, START

temperature and fuel type will vary the

FUEL ON/OFF switch, two INT FUEL TRANS PUMP or

weight of the fuel remaining and the

INT AUX FUEL switches, and GOV AUTO/EMER switch.

actual time the aircraft engine may

The switch over to emergency mode is accomplished by

operate.

Differences

fuel

retarding the throttle to idle or off position and positioning

consumption rates, aircraft attitude and

the GOV AUTO/EMER switch to the EMER position. In

operational condition of the fuel

the EMER position fuel is manually metered to the

subsystem will also affect actual time the

engine, with no automatic control features, by rotating

aircraft engine may operate.

the throttle twist grip.

f Fuel Boost Pump Caution Lights. The LEFT

b. Fuel Quantity Indicator.

The fuel quantity

FUEL BOOST and RIGHT FUEL BOOST caution Lights

indicator is located in the upper center area of the

will illuminate when the left/right fuel boost pumps fail to

instrument panel (fig 2-4). This instrument is a

pump fuel.

The circuits receive power from the 28

Vdc

Change 17 2-7

TM 55-1520-210-10

essential bus. Circuit protection is provided by the

a. Internal Fuel Transfer Switches. Two switches

CAUTION LIGHTS, RH FUEL BOOST PUMP and LH

marked INT AUX FUEL LEFT/RIGHT are mounted in the

FUEL BOOST PUMP circuit breakers. On helicopters

ENGINE control panel (fig 2-6). Placing the switches to

prior to Serial No.

69-15292 a FUEL TANK SUMP

the forward position energizes the auxiliary fuel system.

PUMP circuit breaker is used instead of RH and LH

Fuel is transferred to the main fuel cells. An overfill limity

BOOST PUMP circuit breakers.

switch Is installed in the main fuel tank to prevent the

auxiliary fuel pumps from overfilling the main fuel cells.

2-27. Auxiliary Fuel System.

Power is supplied by the dc essential bus and protected

by the FUEL TRANS PUMP circuit breaker.

Complete provisions have been made for installing an

auxiliary fuel equipment kit in the helicopter cargo

b. Auxiliary Fuel Low Caution Light. An AUX FUEL

passenger compartment. Two crashworthy bladder type

LOW caution light Is provided to indicate when the

tanks can be installed on the aft bulkhead and

auxiliary fuel tanks are empty. The light will Illuminate

transmission support structure.

This allows the

only when the fuel transfer switches are in the forward

helicopter to be serviced with an additional 300 U.S.

position, and the auxiliary tanks are empty. The circuit

gallons of fuel (table 2-1).

receives power from the 28 Vdc essential bus and is

protected by the CAUTION LIGHTS circuit breaker.

Section V. FLIGHT CONTROL SYSTEM

2-28. Description. The flight control system is a

perpendicular to the mast Is controlled by the hydraulic

hydraulic assisted positive mechanical type, actuated by

dampers. By adjusting the dampers, positive dynamic

conventional helicopter controls. Complete controls are

stability can be achieved, and still allow the pilot

provided for both pilot and copilot. The system includes

complete responsive control of the helicopter.

a cyclic system, collective control system, tail rotor

system, force the system, synchronized elevator, and a

2-30. Collective Control System. The collective pitch

stabilizer bar.

control lever controls vertical flight (fig 2-5). When the

lever is in full down position, the main rotor is at

2-29. Cyclic Control System. The system is operated

minimum pitch. When the lever is in the full up position,

by the cyclic stick movement (fig 2-5). Moving the stick

the main rotor Is at maximum pitch. The amount of lever

in any direction will produce a corresponding movement

movement determines the angle of attack and lift

of the helicopter which Is a result of a change in the

developed by the main rotor, and results in ascent or

plane of rotation of the main rotor. The pilot cyclic

descent of the helicopter. Desired operating friction tan

contains the cargo release switch, radio ICS switch,

be induced into the control lever by hand-tightening the

armament fire control switch, hoist switch and force the

friction adjuster (fig 2-5). A grip-type throttle and a switch

switch. Desired operating friction can be induced into the

box assembly are located on the upper end of the

control stick by hand tightening the friction adjuster.

collective pitch control lever. The pilot switch box

contains the starter switch, governor rpm switch, engine

a. Synchronized Elevator.

The synchronized

idle stop release switch, and landing light/searchlight

elevator (fig 2-1) is located on the tailbroom. It is

switches. A collective lever down lock is located on the

connected by control tubes and mechanical linkage to

floor below the collective lever. The copilot collective

the fore and-aft cyclic system. Fore and aft movement of

lever contains only the grip-type throttle, governor rpm

the cyclic control stick will produce a change in the

switch, and starter switch when installed. The collective

synchronized elevator attitude.

This improves

pitch control system has built-in breakaway (friction)

controllability within the cg range.

force to move the stick up from the neutral (center of

travel) position of eight to ten pounds with hydraulic

b. Stabilizer Bar. The stabilizer bar is mounted on

boost ON.

the main rotor hub trunnion assembly in a parallel plane,

above and at 90 degrees to the main rotor blades. The

2-31. Tail Rotor Control System. The system is

gyroscopic and inertial effect of the stabilizer bar will

operated by pilot/copilot anti-torque pedals (fig 2-5).

produce a damping force in the rotor rotating control

Pushing a pedal will change the pitch of the tail rotor

system and thus the rotor.

When an angular

blades resulting in directional control. Pedal adjusters

displacement of the helicopter/mast occurs the bar tends

are provided to adjust the pedal distance for individual

to remain in its trim plane. The rate at which the bar

comfort. A force the system is connected to the

rotational plane tends to return to a position

directional controls.

2-8

Change 17

TM 55-1520-210-10

2-32. Force Trim System. Force centering devices are

control stick and anti-torque pedals. A FORCE TRIM

incorporated in the cyclic controls and directional pedal

ON/OFF switch is installed on the miscellaneous control

controls. These devices are installed between the cyclic

panel to turn the system on or off (Fig 2-6). These

stick and the hydraulic servo cylinders, and between the

forces can be reduced to zero by pressing and holding

anti-torque pedals and the hydraulic servo cylinder. The

the force trim push-button switch on the cyclic stick grip

devices furnish a force gradient or "feel’ to the cyclic

or moving the force trim switch to OFF.

Section VL HYDRAULIC SYSTEM

2-33. Description. The hydraulic system is used to

feedback to the cyclic and collective in the event of

minimize the force required by the pilot to move the

hydraulic system malfunction.

cyclic, collective and pedal controls. A hydraulic pump,

mounted on and driven by the transmission supplies

2-34. Control Switch. The hydraulic control switch is

pressure to the hydraulic servos. The hydraulic servos

located on the miscellaneous panel (Fig 2-6). The switch

are connected into the mechanical linkage of the

is a two-position toggle type labeled HYD CONTROL

helicopter flight control system.

Movement of the

ON/OFF. When the switch is in the ON position,

controls in any direction causes a valve, in the

pressure is supplied to the servo system. When switch

appropriate system, to open and admit hydraulic

is in the OFF position the solenoid valve is closed and no

pressure which actuates the cylinder, thereby reducing

pressure is supplied to the system. The switch is a fail-

the force-load required for control movement.

safe type. Electrical power is required to turn the switch

Irreversible valves are installed on the cyclic and

off.

collective hydraulic servo cylinders to prevent main rotor

Change 17

2-8.1/(2-8.2 blank)

TM 55-1520-210-10

2-35. Reservoir and Sight Glass. The hydraulic

provided to permit ready visual access to the filter

reservoir is a gravity feed type and is located at the right

indicator. The transparent window is located on forward

aft edge of the cabin roof (fig 2-10). The reservoir and

face of the transmission bulkhead.

sight gage are visible for inspection through a plastic

window in the transmission fairing.

2-37.

Hydraulic Pressure Caution Light. Low

hydraulic system pressure will be indicated by the

2-36. Hydraulic Filter. A line filter is installed on

illumination of

helicopters prior to Serial No. 68-16050. Thus filter has

HYD PRESSURE

segment on

the

no indicator. Helicopters Serial No

63-16050 and

CAUTION

subsequent, or those modified by MWO have an

panel. Moderate feedback forces will be noticed in the

improved filter system. When the filter is clogged it will

controls when moved.

give a visual warning by raising a red indicator button.

The red button pops out when a set differential pressure

2-38. Electrical Circuit Electrical power for hydraulic

across the element is exceeded. Once actuated, the

system control is supplied by the 28 Vdc essential bus.

indicator will remain extended until reset manually When

The circuit Is protected by the HYD CONT circuit

the indicator TM 55-1520-210-10 is in reset position, it

breaker.

will be hidden from view. An inspection window may be

Section VII. POWER TRAIN SYSTEM

2-39. Transmission. The transmission is mounted

A bypass condition will be indicated by extension of a red

forward of the engine and coupled to the power turbine

indicator on the filter head. On Helicopters equipped

shaft at the cool end of the engine by the main

with ODDS, the external oil filter is rated to 3 microns.

dnveshaft. The transmission is basically a reduction

gearbox, used to transmit engine power at a reduced

2-40. Gearboxes.

rpm to the rotor system. A freewheeling unit is

incorporated m the transmission to provide a quick-

a Intermediate Gearbox 42 Degree.

The

disconnect from the engine if a power failure occurs.

degree gearbox is located at the base of the vertical fin.

This permits the main rotor and tail rotor to rotate m

It provides 42 degree change of direction of the tail rotor

order to accomplish a safe auto-rotational landing. The

dnveshaft The gearbox has a self-contained wet sump oil

tail rotor drive is on the lower aft section of the

system. An oil level sight glass, filler cap, vent (fig 2-10)

transmission. Power is transmitted to the tail rotor

and magnetic chip detector are provided.

through a series of driveshafts and gearboxes. The rotor

tachometer generator, hydraulic pump, and main dc

generator are mounted on and driven by the

b. Tail Rotor Gearbox 90 Degree. The 90 degree

transmission. A self-contained pressure oil system is

gearbox is located at the top of the vertical fin. It

incorporated in the transmission. The oil is cooled by an

provides a 90 degree change of direction and gear

oil cooler and turbine fan. The engine and transmission

reduction of the tail rotor driveshaft. The gearbox has a

oil coolers use the same fan. The oil system has a

self-contained wet sump oil system. An oil level sight

thermal bypass capability An oil level sight glass, filler

glass, vented filler cap (fig 2-10) and magnetic chip

cap. and magnetic chip detector are provided. A

detector are provided.

transmission oil filter is mounted in a pocket in the upper

right aft corner of sump case, with inlet and outlet ports

2-41. Driveshafts.

through internal passages. The filter incorporates a

bypass valve for continued oil flow if screens become

A Main Dnveshaft The main dnveshaft connects

clogged. The transmission external oil filter is located in

the engine output shaft to the transmission input drive

the cargo-sling compartment on right side wall, and is

quill.

connected into the external oil line. On Helicopters

equipped with ODDS, a full flow debris monitor with

b. Tall Rotor Driveshaft The tail rotor dnveshaft

integral chip detector replaces the integral oil filter. A

consists of six dnveshaft and four hanger bearing

bypass valve is incorporated, set to open at a set

assemblies. The assemblies and the 42 degree and 90

differential pressure to assure oil flow if filter element

degree gearboxes connect the transmission tail rotor

should become clogged.

drive quill to the tail rotor.

Change 17

2-9

TM 55-1520-210-10

2-42. Indicators and Caution Lights.

a. Transmission Oil Pressure Indicator.

The

breaker. This is a wet bulb system dependent on fluid

TRANS OIL pressure indicator is located in the center

for valid indication.

area of the instrument panel (fig 24). It displays the

transmission oil pressure in psi. Electrical power for the

circuit is supplied from the 28 Vac bus and is protected

e. Transmission and Gearbox Chip Detector

by the XMSN circuit breaker in the ac circuit breaker

panel.

(I) Chip Detector Caution Light. Magnetic

inserts are installed in the drain plugs of the transmission

b. Transmission Oil Pressure Low Caution Light.

sump, 42 degree gearbox and the 90 degree gearbox.

The XMSN OIL PRESS segment in the CAUTION panel

On helicopters equipped with ODDS, the transmission

will illuminate when the transmission oil pressure drops

chip gap is integral to a full-flow debris monitor. When

below about 30 psi.

The circuit receives power from

sufficient metal particles collect on the plugs to close the

the essential bus. Circuit protection is supplied by the

electrical circuit the CHIP DETECTOR segment In the

CAUTION LIGHTS circuit breaker.

CAUTION panel will illuminate. A self-closing, spring-

loaded valve in chip detectors permits the magnetic

c. Transmission Oil Temperature Indicator. The

probes to be removed without the loss of oil The circuit is

transmission oil temperature indicator is located in the

center area of the instrument panel (Fig

24).

The

CAUTION LIGHTS circuit breaker.

indicator displays the temperature of the transmission oil

in degrees Celsius. The electrical circuit receives power

(2) Chip Detector Switch. A CHIP DET switch

from the essential bus and is protected by the TEMP IND

(fig 2-6) is installed on a pedestal mounted panel. The

ENG XMSN circuit breaker in the dc breaker panel. This

switch is labeled BOTH, XMSN, and TAIL ROTOR and is

is a wet bulb system dependent on fluid for valid

spring-loaded to the BOTH position. When the CHIP

indication.

DETECTOR segment in the CAUTION panel lights up

position the switch to XMSN, then TAIL ROTOR to

d. Transmission Oil Hot Caution Light. The XMSN

determine the trouble area. CHIP DET caution light will

OIL HOT segment in the CAUTION panel will illuminate

remain on when a contaminated component is selected.

when the transmission oil temperature is above IIO’C

The light will go out if the noncontaminated component is

(230’F). The circuit receives power from the essential

selected.

bus and is protected by the CAUTION LIGHTS circuit

Section VIII. ROTORS

243. Main Rotor.

attached to a support affixed to the mast. The hub

springs 2-10

Change 17 provide an additional margin

a. Description. The main rotor is a two bladed

of safety in the event of on inadvertent excursion of the

senm-rigid, seesaw type. The two types of rotor blades

helicopter beyond the approved flight envelope.

are metal and composite material and must not be

intermixed. The two blades are connected to a common

b. RPM Indictor. The rpm indicator is part of the

yoke by blade grips and pitch change bearing with

dual tachometer (fig 24). The tachometer inner scale

tension straps to carry centrifugal forces. The rotor

displays the rotor rpm. The inner scale pointers is

assembly is connected to the mast with a nut. The nut

marked with an ’R’.

has provisions for hoisting the helicopter. A stabilizer bar

is mounted on the trunnion 90 degrees to the main rotor.

2-44. Tail Rotor. The tall rotor is a two-bladed semi-

Blade pitch change is accomplished by movements of

rigid delta-hinge type. Each blade is connected to a

the collective and cyclic controls. The main rotor is

common yoke by a grip and pitch change bearings. The

driven by the transmission through the mast. The mast

hub and blade assembly is mounted on the tail rotor

is tilted 5 degrees forward.

shaft with a delta-hinge trunnion and a static stop to

minimize rotor flapping.

Blade pitch change is

al. Hub Spring. As an aid in controlling rotor

accomplished by movement of the anti-torque pedals

flapping a hub spring kit has been installed in the rotor

which are connected to a pitch control system through

system for those helicopters modified by MWO 55-1520-

the tail rotor (90 degree) gearbox. Blade pitch change

242-50-1. Two nonlinear elastometric springs are

servos to offset torque and provide heading control.

2-10 Change 17

TM 55-1520-210-10

Section DC UTILITY SYSTEMS

2-45.

Pilot Heater. The piton tube is equipped with an

2-50. External Cargo Rear View Mirror. A mirror may

electrical heater (Figure 2-1). The PITOT HTR switch is

be installed under the right lower nose window to give the

on the overhead console panel (Figure 2-5). ON position

pilot clear visibility of the external cargo. This mirror may

activates the heater in the tube and prevents ice from

be removed and stowed in the heater compartment when

forming m the pilot tube. OFF position de-activates the

provisions are installed.

heater. The electrical circuit for the system receives

power from the essential bus and is protected by the

2-51. Windshield Wiper.

PITOT TUBE HTR circuit breaker.

2-46. Heated Blanket Receptacles. Two or six

electrical receptacles are provided to supply 28 Vdc for

heated blankets. They are mounted on the inside cabin

Do not operate the wiper on a dry or

roof structure aligned with the forward edge of the

dirty windshield.

transmission support structure. The electrical circuit for

the receptacles receive power from the nonessential bus.

a Two windshield wipers are provided, one for the

Circuit protection is provided by the HEATED BLANKET

right section of the windshield and one for the left section

circuit breakers.

of the windshield.

2-47. Data Case. A data case for maps, flight reports,

b. The wipers are driven by electric motors with

electric power supplied by the dc electrical system.

etc., has been provided and is located at the aft end of

Circuit protection is provided by WINDSHIELD WIPER

the pedestal (Figure 2-5).

PILOT and WINDSHIELD WIPER COPILOT circuit

breakers on the dc circuit breaker panel.

2-48. Blackout Curtains. Provisions have been made

for Installing blackout curtains behind pilot and copilot

c The windshield wiper switches on the overhead

seats and between forward and aft cabin sections. Other

console mounted MISC panel (Figure 2-5) have five

blackout curtains may be installed over both cabin door

positions: HIGH, MED, LOW, OFF, and PARK.

windows and window in removable doorpost.

d The panel also has a selector which permits the

2-49. Blood Bottle Hangers. Provisions have been

operation of windshield wiper for pilot, copilot or both as

made for six blood bottle hangers on the inside of the

desire.

cabin roof structure within easy reach of the medical

attendant station, for administration of blood to litter

patients in flight.

Section X. HEATING AND VENTILATION

2-52. Ventilating System.

these helicopters. They are the bleed air heater,

combustion heater, and the auxiliary exhaust heat

a Description The ventilating system consists of

exchanger. Each system is described separately in the

four independently controlled exterior air scoop

following paragraphs.

ventilators.

Two single orifice air scoops are located

on * top of the cabin section, and two double orifice air

a Bleed Air Heating and Defrosting System There

scoops are on top of cabin. The amount of air entering

are some differences in the bleed air heating systems in

the cabin through the ventilators is regulated by the

use.These differences are shown in Figure 2-7 with the

butterfly valve control.

following exception: helicopters prior to Serial No. 5-

9565 have under-seat heater outlets; subsequent

b. Operation Rotate butterfly valve control to

helicopters have aft pedestal outlets instead of the

desired position to provide outside air for flight.

under-seat outlets. Heat is supplied to all bleed air

heaters

2-53. Heating and Defrosting System. Three different

types of heating and defrosting systems may be used on

Change 17

2-11

TM 55-1520-210-10

by the compressor bleed air system. Electric power for

operating after shutdown to prevent residual heat

operation of the controls Is supplied from the essential

buildup. If blower air pressure drops too low, the

bus and is protected by the CABIN HEATER CONT

combustion heater will stop automatically. An overheat

circuit breaker. On helicopters Serial No.

66-16868

switch also automatically turns the heater off in the event

through

70-16518, temperature is controlled by a

of malfunction. The starting cycle has to be repeated to

thermostat located on the right doorpost. Helicopters

start the combustion heater. Electric power to operate

Serial No. 71-20000 and subsequent are protected by

the heater controls is supplied from the essential bus and

two circuit breakers marked CABIN HEATER OUTLET

is protected by the CABIN HEATER CONT circuit

VALVE and CABIN HEATER AIR VALVE. Refer to

breaker. Refer to Figure 2-7 for controls and their

Figure 2-7 for controls and their function.

function.

b Combustion Heating and Defrosting System

c Auxiilary Exhaust Heater System The auxiliary

With the combustion heater installed, a combination of

exhaust heater system consists of an exhaust gas

bleed air heat and combustion heat is available for

exchanger, and a bleed air driven fan for circulating

heating, Bleed air may be used for defrosting and

ambient air through the heat exchanger. A mixing valve

combustion heat for heating, or combustion heat may be

controls air to maintain the desired outlet temperature.

used for defrosting only with bleed air heat off. The

The exhaust heater system controls consist of the cabin

MAIN FUEL switch must be ON, actuating the right boost

heating panel (Figure 2-7), a thermostat dial on the right

pump, before fuel is available for combustion heater

door post and the air directing lever on the pedestal.

operation (Figure 2-6). A purge switch keeps the blowers

Section XI. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM

2-54. DC and AC Power Distribution. Figure 2-8

2-57. Battery.

depicts the general schematic of the dc and ac power

distribution system. The dc power is supplied by the

battery, main generator, standby starter-generator, or the

external power receptacle. The 115 Vac power is

supplied by the main or spare inverters. The 28 Vac

If battery overheats, do not open

power Is supplied by a transformer which is powered by

battery compartment. Battery fluid

the inverter.

will cause burns. An overheated

battery may cause thermal burns and

2-55. DC Power Supply System. The dc power supply

may explode.

system is a single conductor system with the negative

leads of the generator grounded in the helicopter

The battery supplies approximately 2-4Vdc power to the

fuselage structure. The main generator voltage will vary

power distribution system when the generators and

from 27 to 28.5 depending on the average ambient

external power receptacle are not in operation (Figure 2-

temperature. In the event of a generator failure-the

1).

nonessential bus is automatically de-energized. The

pilot may override the automatic action by positioning the

2-58. Main and Standby Starter-Generator. The 30

NON-ESS BUS switch on the DC POWER control panel

volt 300 ampere main generator is mounted on and

to MANUAL ON.

driven by the transmission. A standby starter-generator,

rated at

300 amperes is mounted on the engine

2-56. External Power Receptacle. The external power

accessory drive section.

The standby furnishes

receptacle (Figure 2-1) transmits the ground power unit

generator power in

28 Vdc power to the power distribution system. A 7.5

KW GPU is recommended for external starts.

the event of main generator failure.

2-12 Change 17

TM 55-1520-210-10

2-59. DC Power Indicators and Controls.

bus to receive dc power from the main generator. The

a. Main Generator Switch The MAIN GEN switch

TM 55-1520-210-10 MANUAL ON position permits the

(fig 2-5) is on the overhead console DC POWER panel.

nonessential bus to receive power from the standby

In the ON position the main generator supplies power to

generator when the main generator is off line.

the distribution system. The RESET position is spring-

loaded to the OFF position. Momentarily holding the

e. DC Voltmeter Selector Switch The VM switch is

switch to RESET position will reset the main generator.

located on the DC POWER control panel (fig 2-5).

The OFF position isolates the generator from the

The switch permits monitoring of voltage being delivered

system. The circuit is protected by the GEN & BUS

from any of the following; BAT, MAIN GEN, STBY GEN,

RESET in the dc circuit breaker panel.

ESS BUS, and NON-ESS BUS.

b. Battery Switch The BAT switch is located on the

f. DC Voltmeter. The dc voltmeter is located in

DC POWER control panel (fig 2-5). ON position permits

the center area of the instrument panel and is labeled

the battery to supply power. ON position also permits

VOLT DC (fig 2-4). Direct current voltage is indicated on

the battery to be charged by the generator. The OFF

the voltmeter as selected by the VM switch in the

position isolates the battery from the system.

overhead console.

c. Starter-Generator Switch The STARTER GEN

g. DC Loadmeters-Main and Standby. Two direct

switch is located on the DC POWER control panel (fig 2-

current loadmeters are mounted in the lower center area

5).The START position permits the starter-generator to

of the instrument panel (fig

24).

The MAIN GEN

function as a starter. The STBY GEN position permits

loadmeter indicates the percentage of main generator

the starter-generator to function as a generator.

rated capacity being used. The STBY GEN loadmeter

indicates the percentage of standby generator rated

d. Nonessential Bus Switch. The NON-ESS BUS

capacity being used. The loadmeters will not indicate

switch is located on the DC POWER control panel (fig 2-

percentage when the generators are not operating.

5). The NORMAL ON position permits the nonessential

Change 17

2-12.1/(2-12.2 blank)

TM 55-1520-210-10

2-60. DC Circuit Breaker Panel. The dc circuit breaker

TM 55-1520-210-10 bus. Circuit protection is provided

panel Is located in the overhead console (fig 2-5). In the

by the INVTR CONT circuit breaker.

"pushed in" position the circuit breakers provide circuit

protection for dc equipment. In the "pulled out’ position

b. AC Failure Caution Light. The INST INVERTER

the circuit breakers deenergize the circuit. In the event

caution light will illuminate when the inverter m use fails

of an overload the circuit breaker protecting that circuit

or when the INVTR switch is m the OFF position.

will "pop out". Each breaker is labeled for the particular

circuit it protects. Each applicable breaker is listed m

the paragraph descanting the equipment it protects.

c AC Voltmeter Selector Switch. The

PHASE VM switch is located on the AC POWER control

2-61. AC Power Supply System. Alternating current is

panel (fig 2-5). The switch is used to select any one of

supplied by two inverters (fig 2-8). They receive power

the three phases of the 115 Vac three-phase current for

from the essential bus and are controlled from the AC

monitoring on the ac voltmeter. The three positions on

POWER control panel (fig 2-5).

the switch are: AB, AC, and BC. Each position indicates

that respective phase of the 115 Vac on the ac voltmeter

24-62. Inverters. Either the main or spare inverter (at

the pilots option) will supply the necessary 115 Vac to the

d. AC Voltmeter. The ac voltmeter Is mounted on

distribution system. The inverters also supply 115 Vac to

center area of the instrument panel (fig 24). The ac

I the 28 volt ac transformer which in turn supplies 28 Vac

voltage output from the inverter (main or spare) is

to the necessary equipment. Circuit protection for the

indicated on this instrument. The voltage indicated on

inverters is provided by the MAIN [NVTR PWR and

any of the three selected positions should be 1 12 to 1 18

SPARE NVTR PWR circuit breakers.

Vac 2-64. AC Circuit Breaker Panel. The ac circuit

breaker panel is located on the right side of the pedestal

2-63. AC Power Indicators and Controls.

panel (fig 2-5). The circuit breakers in the "pushed m"

position provide circuit protection for the equipment. The

a Inverter Switch. The INVTR switch is located on

breakers m the "pulled out" position de-energize the

the AC POWER control panel in the overhead console

circuit. The breakers will pop out automatically m the

(fig 2-5). The switch is normally in the MAIN ON

event of a circuit overload. Each breaker is labeled for

position, to energize the main inverter. In the event of a

the particular circuit it protects. Each applicable breaker

main inverter failure the switch can be positioned to

Is listed in the paragraph describing the equipment it

SPARE ON to energize the spare inverter. Electrical

protects.

power to the INVTR switch is supplied from the essential

Section XII. LIGHTING

NOTE

(1) Configurations.

Two red lights are

Visible light means the light is visible to

mounted on the left side of the fuselage, one above and

the unaided eye. NVG light means the

one below the cabin door. Two green lights are mounted

light is visible only with the aid of the

on the right side of the fuselage, one above and one

Night Vision Goggles.

below the cabin door. Two white lights are mounted on

top of the fuselage just inboard of the red and green

2-65. Position Lights. The position lights consist of

lights. One white light is mounted on the bottom center

eight visible lights and five NVG lights (fig 2-1).

of the fuselage, and one white light is mounted on the

a Visible Position Light Lights.

Change 17

2-13

TM 55-1520-210-10

tailboom vertical fin. Electric power to operate the lights

b. Operation The ON position of the ANTI COLL

is supplied from the essential bus. Circuit protection is

light switch illuminates the anti-collision light and starts

provided by the NAV LIGHTS circuit breaker in the dc

rotation of the light (fig 2-5). OFF position de-energizes

circuit breaker panel. Some position lights may be

the light.

protected by the FUS LIGHTS circuit breakers.

NOTE

(2) Operation of Visible Position Lights. The

The IR Band-Pass Filter (I.R. COVER)

position lights are controlled by the POSITION switches

is authorized to be installed on either the

on the EXT LTS panel on the overhead console (fig 2-5).

landing light or the searchlight or may be

A three-position switch permits selection of STEADY,

removed from the aircraft for unaided

OFF, or FLASH. Another two-position switch control

(NON-NVG) flights dependent

brilliance and is marked DIM and BRIGHT. When the

operational

considerations

three-position switch is in STEADY position, all eight

requirements. Both the landing light and

navigation lights are llumunated. In FLASH position, on

searchlight operation with the band-pass

helicopters prior to Serial No.

64-13901, the colored

filter installed is the same as normal

lights illuminate alternately with the white lights. On later

landing

light

and

searchlight

models only the colored lights and the aft white light

operation, except that NVGS must be

flash.

worn to use it. The band-pass filter (P/N

EGD-0931 1) shall not be used with

b. NVG Poisson Lights

lamps exceeding 250 watts.

(I) Configuration. These lights are invisible to

2-67. Landing Light.

the unaided eye They are designed to provide observed

aircraft position, attitude, and distance during covert

a General The landing light is flush-mounted to the

formation NVG flight and other covert multi-aircraft NVG

underside of the fuselage (fig 2-1). It may be extended or

operations. Lights are located on the top left and right

retracted to improve forward illumination. Electric power

side above the jump door of each side and one each

to operate the system is supplied from the essential bus.

under the pilot and copilot doors. The rear NVG position

Circuit protection is provided by the LDG LIGHT PWR

light is located on a mount under the visible position light.

and LDG SEARCH LIGHT CONT circuit breakers.

(2) Operation of the NYG Position Lights. The

b. Operation Landing light switches are on the pilot

control panel for the lights is located in the frontmost

collective lever switch box (fig 2-5). The ON position of

panel of the left overhead console (fig

2-5). Five

the LDG LT switch causes the landing light to illuminate;

intensity positions are provided on the control panel;

OFF turns the light off. The EXT position of the LDG LT

OFF, 1, 2, 3, 4, and BRT (BRIGHT). The lights are

EXT OFF RETR extend the landing light to the desired

invisible to the unaided eye and should be checked or

position; RETR position retracts the light. The OFF

otherwise viewed with AN/PVS-5 AN/AVS-6 or AN/PVS-7

position stops the light during extension or retraction.

NVG. The visible position EXT LTS must be in the off

The light automatically stops at the full extend/retract

position when the NVG lights are used. The NVG lights

position.

should be in the off position when not being used with

NVG. The NVG lights do not flash. Circuit protection is

provided by the NAV lights circuit breaker.

2-66. Anti-Collision Light.

a General. The anti-collision light is located on

the top aft fuselage area (fig 2-I). Electric power to

operate the light is supplied from the essential bus.

Circuit protection is provided by the ANTI COLL LIGHT

circuit

breaker.

2-14 Change 17

TM 55-1520-210-10

2-68. Searchlight.

b. Operation To operate the FWD dome light

position the FWD switch to WHITE for white light NVG

a. General The searchlight is flush-mounted to the

for green light and OFF to turn light off. The aft dome

underside of the fuselage (fig 2-1). The light can be

lights panel has two switches. The WHITEIOFF NVG

extended and retracted for search illumination. At any

switch functions are the same as the FWD Switch.

desired position in the extend or retract arc, the light may

Rotation of the rheostat marked OFF/MED/BRT

be stopped and rotated to the left or right. Electric power

increases or decreases the brightness of the aft dome

to operate the light is supplied from the essential bus.

lights.

Circuit protection is provided by the SEARCHLIGHT

PWR and LDG & SEARCHLIGHT CONT circuit

2-70. Cockpit Map Lights.

breakers.

a General. Two cockpit lights (NVG Green) are

b Operation. The pilot SL switch ON position

provided, one above the pilot and one above the copilot

illuminates the light (fig

2-5). The OFF position

(fig 2-5). Each light is controlled individually The lights

deactivates the light. The STOW position retracts the

receive power from the essential bus and are protected

light into the fuselage well.

by the COCKPIT LIGHTS circuit breaker.

2-69. Dome Lights.

b. Operation Rheostat switches are part of each

light assembly. Brightness is increased by turning the

NOTE

rheostat

clockwise

dimmed

turning

Aircraft modified by MWO

55-1520-210-5012 the

counterclockwise. These lights are for NVG or unaided

dome light switch, if installed, is not connected.

eye use.

a General The dome lights provide overhead

2-71. Instrument Lights. The instrument lights control

lighting for the cabin area The forward light is controlled

panel is located in the overhead console (fig 2-5). The

by the FWD switch on the DOME LT panel when

panel contains six switch/rheostats for activating and

installed on the overhead console.

When

the

controlling the brightness of the various instrument lights.

DOME LT panel is not installed the FWD and AFT

Each switch/rheostat functions the same. The OFF

DOME LT are controlled by the aft switch in the roof.

position de-energizes the circuit, clockwise rotation

The aft dome lights are controlled by the switch on the

increases brightness of the lights and counterclockwise

AFT DOME LTS panel on the roof. Electric power to

rotation decreases brightness.

The lights of all

operate the dome light is supplied from the 28 Vdc

instruments receive electric power from the 28 Vdc

essential bus. Circuit protection is provided by the

essential bus, except the pilot attitude indicator, the pilot

DOME LIGHTS circuit breaker.

RMI, and the turn and slip indicators which receives 5

Vdc from the essential bus through resistor R24. On

UH-IH helicopters equipped I with radar altimeters, the

pilot attitude indicator, lot’s RMI, and the turn and slip

indicators receive the 5 Vdc from the pilot solid state

device for providing 5 Vdc from the 28 Vdc essential bus.

On UHI H helicopters equipped with radar altimeters,

instruments are illuminated by 28 Vdc from the 28 Vdc

essential bus. On UHI H aircraft the instrument lighting

is protected by CONSOLE PEDLIGHTS, INST PANEL

LIGHTS, and INST SEC LIGHTS circuit breakers. On

UHI H helicopters equipped I with radar altimeters, the

measurements lights are protected by the same circuit

breakers as the UHI H aircraft, except the PILOT

VOLT LIGHTS circuit breaker is included in the

protection. On UH-IH helicopters equipped with radar

altimeters, the "HI" and "LO" indicators and the digital

readout illumination levels on the radar altimeter are

controlled by the pilot’s and co-pilot’s lighting rheostats.

Change 17

2-15

TM 55-1520-210-10

a. Pilot Instrument Lights The pilot instrument lights

furnish illumination for the following instruments; gas

producer tachometer, torquemeter, exhaust temperature

(1) The four secondary instrument lights are (

indicator, dual tachometer, airspeed

indicator, clock,

spaced across the top of the instrument panel shield (fig

vertical velocity indicator, turn

and slip

indicator,

24). These lights furnish secondary illumination for the

altimeter, attitude indicator, radio magnetic indicator,

instrument panel face. The lights are activated and

standby compass, pilot collective switch box and radar

controlled by the switch/rheostat marked SEC on the

altimeter (if installed). These lights are all on one circuit

INST LTG control panel. Circuit protection is provided by

and are controlled by the switch/rheostat marked PILOT

the INST SEC LIGHTS circuit breaker.

on the INST LTG control panel. DIM switch controls 23,

page 3-33. Circuit protection is provided by the INST

(2) Pedestal Utility Light. The pedestal utility

PANEL LIGHTS and PILOT 5 VOLT LIGHT (UH-IV)

light is provided for general use but also provides

circuit breakers.

illumination for radios which are not illuminated. The

light is operated by an OFF and ON switch on the panel

b. On UH-IV aircraft when the radar altimeter,

on which it is mounted. The intensity of the Light is

AN/APNV-209 is installed, turning the pilot rheostat to

controllable from full bright to dim by a rheostat on the

OFF I provides full illumination to the digital readout and

some panel. The light Is protected by the COCKPIT LTS

HI-LO warning lights on the pilot and copilot height

circuit breaker.

indicators. This feature enables the pilot and copilot to

read the displays during daytime operations.

f Emergency Llghning.

Master Caution Fire

Warning and Low RPM Indicators. The Master Caution,

c Copilot Instrument Lights.

The

copilot

Fire, and Low RPM Warning Indicators are equipped

instrument lights furnish illumination for the instruments

with flip filters for use during NVG operations. The

on the copilot section of the instrument panel. These

indicators must be uncovered for unaided operations

instruments consist of an airspeed indicator, attitude

during daylight and night.

indicator, radar altimeter (if installed on UH-1V),

altimeter, vertical velocity indicator and radio magnetic

indicator. The copilot instrument lights are all on one

2-72.

Overhead Console Panel Lights. The

circuit, and are controlled by the switch/rheostat marked

overhead console panel lights furnish illumination for all

COPILOT on the INST LTG control panel. Circuit

overhead panels (fig 2-5). Each panel Is individually

protection is provided by INST PANEL LIGHTS circuit

illuminated and control is accomplished by the

breaker. Circuit protection is provided by INST PANEL

switch/rheostat, marked CONSOLE on the INST LTG

LIGHTS and COPILOTS

5 VOLT LIGHTS circuit

control panel. Circuit protection is provided by the

breaker.

CONSOLE PED LIGHTS circuit breaker.

d. Engine Instrument Lights.

The engine

NOTE

instrument lights furnish illumination for the following

All "press-to-test" and cateye indicators

instruments: transmission oil temperature, fuel quantity,

are dimming type NVG green.

transmission oil pressure, engine oil pressure,

Differentiation as to which indicator is

loadmeters, ac voltmeter, fuel pressure indicator, engine

illuminated must be determined by

oil temperature gage, and dc voltmeter Each instruments

location since color coding is not used.

is individually illuminated and control is accomplished by

In addition, the operator must return all

the switch/rheostat marked ENGINE on the INST LTG

indicators to the full bright position to

control panel. Circuit protection is provided by the INST

assure visibility.

On applicable

PANEL LIGHTS circuit breaker

indicators,

dimming

may

accomplished during

e Secondary Instrument Lights.

pre-flight check to meet the operators needs.

2-16

Change 17

TM 55-1520-210-10

2-73. Pedestal Lights. The pedestal lights provide

transmission oil sight gage. The circuit is activated by

illumination for the control panels on the pedestal (fig 2-

TM 55-1520-210-10 a button-type switch marked XMSN

5). Most panels are individually illuminated and intensity

OIL LEVEL LT SWITCH and is located on the right side

control is accomplished by the switch/rheostat marked

of the transmission forward bulkhead. Electric power for

PED on the INST LTG control panel overhead. On some

the transmission oil level light circuit is supplied by the

panels, the internal lighting has been discontinued.

battery. Circuit protection is provided by the battery

These panels are illuminated by the goosenecked utility

voltmeter circuit breaker located in the oil cooler

light mounted to the rear of the pedestal.

The

compartment or forward radio compartment.

illumination can be placed on any area of the pedestal by

adjusting the gooseneck position. The intensity and ON-

2-75. Spare Lamp Kit. The spare lamp kit is located on

OFF function is controlled by the switch/rheostat

the left side of the overhead console. The kit contains

mounted on the panel with the utility light. Circuit

spare light bulb for the segment panel lights, the

protection for the pedestal lighting is provided by the

instrument lights, pedestal and overhead console lights,

CONSOLE PED LIGHTS circuit breaker.

master caution and segment caution lights, all press-to-

test lights, the rpm and fire warning lights, and the dome

2-74. Transmission Oil Level Light. A transmission oil

lights. All bulbs except the dome light bulbs may be

level light is installed to provide illumination to check the

replaced without the use of tools.

Change 17

2-16.1/(2-16.2 blank)

TM 55-1520-210-10

Section XIIL FLIGHT INSTRUCTIONS

2-76.

Airspeed Indicators. The pilot and copilot

airspeed indicators display the helicopter indicated

b. Copilot Attitude Indicator The copilot attitude

airspeed (IAS) in knots (fig 2-4). The IAS is obtained by

indicator is located in the copilot section of the

measuring the difference between impact air pressure

instrument panel (fig 2-4). It is operated by 115 Vac

from the pitot tube and the static air pressure from the

power supplied by the inverter. Circuit protection is

static ports (fig 2-1).

provided by the COPILOT ATITD circuit breakers in the

ac circuit breaker panel. In a climb or dive exceeding 27

NOTE

degrees of pitch the horizontal bar will stop at the top or

Indicated airspeeds are unreliable below 20

bottom of the case and the sphere then becomes the

knots due to rotor downwash.

reference. The copilot attitude indicator may be caged

manually by pulling the PULL TO CAGE knob smoothly

2-77.

Turn and Slip Indicator. The turn and slip

away from the face of the instrument to the limit of Its

indicator displays the helicopter slip condition, direction

travel and then releasing quickly.

of turn and rate of turn (fig 2-4). The ball displays the

slip condition. The pointer displays the direction and rate

2-81.

Free-Air Temperature Indicator (FAT).The

of the turn. The circuit receives power from the essential

free-air temperature indicator is located at the top center

bus and is protected by the TURN & SLIP IND circuit

area of the windshield (fig 2-5). The indicator displays

breaker.

the free air temperature m degrees Celsius.

2-78.

Vertical Velocity Indicator.

The vertical

2-82.

Standby Compass. The standby (magnetic)

velocity indicator displays the helicopter ascent and

compass is mounted in a bracket at the center right edge

descenated by the rate of atmospheric pressure change.

of the instrument panel (fig 24). A deviation m magnetic

compass indications will occur when the landing light,

2-79.

Pressure Altimeter. The pressure altimeter

searchlights, or pitot heat are turned on.

(ALT) furnishes direct readings of height above sea level

and is actuated by the pitot static system (fig 24). Two

2-83. Fire Detector Warning System. A FIRE

altimeters are provided, one for the pilot and one for the

WARNING light is located in the upper nght section of

copilot. (Refer to chapter 3 for operation.)

the instrument panel (fig 24). The press to test (FIRE

DETECTOR TEST) test switch is located to the left of

2-80.

Attitude Indicators.

the fire warning light. Excessive heat m the engine

a. Pilot Attitude Indicator. The pilot attitude indicator

compartment causes the FIRE light to illuminate.

is located on the pilot section of the instrument panel (fig

Pressing the press-to-test switch also causes the light to

24). The indicator displays the pitch and roll attitude of

Illuminate. Electric power for the circuit is supplied from

the helicopter. An OFF warning flag in the indicator is

the 28 Vdc essential bus and is protected by the FIRE

exposed when electrical power to the system is

DET circuit breaker.

removed. However, the OFF flag will not indicate

internal system failure. The attitude indicator has an

2-84. Master Caution System.

electrical trim in the roll axis m addition to the standard

NOTE

pitch trim. The attitude indicator is operated by 115 Vac

Aircraft are equipped with NVG compatibility

power, supplied by the inverter. Circuit protection is

devices, flip-filters for the "Master Caution,"

provided by the PILOT ATTD circuit breakers in the ac

"Low RPM," and "Fire Warning" indicators.

circuit breaker panel.

These filters must be flipped over away from

the indicators during visual flight conditions. A

slide drawer filter is also provided for the

caution panel. This filter must be stowed in the

pedestal stowing position when not being used

for NVG flight. To stow, lift the front end of the

The copilot attitude indicator shall be caged only in a

filter to the vertical position and allow the filter to

straight and level attitude. The caging knob shall

gently slide into the vertical cavity in the

never be pulled violently.

pedestal above the caution panel.

Change 17

2-17

TM 55-1520-210-10

a NVG Flight Conditions.

cause the illumination of all the individually worded

segments and the master caution indicator. Only the

(1) Follow all procedures used for visual flight

lamp circuitry is tested; the condition circuitry is not.

conditions, except the "Master Caution," ’Low RPM," and

Testing of the system will not change any particular

’Fire Warning flip-filters and "Caution Panel" slide drawer

combination of fault indications which might exist prior to

filter must be placed over the indicators.

testing. The worded segments will remain illuminated as

(2) Flip instrument panel indicator filters over

long as fault condition or conditions exist, unless the

indicators and press lightly in place to avoid light leakage

segment is rotated.

around edges.

(3) Gently pull the slide drawer filter up from

d. Electrical Power. Electric power for the master

stowed position until it is at the top vertical position and

caution system is supplied from the essential bus.

place it over the caution panel.

Circuit protection is provided by the CAUTION LIGHTS

circuit breakers.

b. Master Caution Indicator The master caution

indicator light on the instrument panel will Illuminate

2-85. RPM High-Low Limit Warning System. The

when fault conditions occur (figure 2-4). This illumination

rpm high-low limit warning system provides the pilot with

alerts the pilot and copilot to check the caution panel for

an immediate warning of high and low rotor or engine

the specific fault condition.

rpm. Main components of the system are a detector

unit, warning light and audio signal circuit, low RPM

c. Caution Panel. The CAUTION panel is located

AUDIO/OFF switch, and electrical wiring and connectors.

on the pilot side of the pedestal (figure 2-9). Worded

The warning light and audio warning signal systems are

segments illuminate to identify specific fault conditions.

activated when any one of the following rpm conditions

The worded segments are readable only when the light

exist:

illuminates. When a light illuminates, flickers or full

illumination, it indicates a fault condition. Refer to figure

a Warning light only:

2-9 for explanation of the fault conditions.

(1) For rotor rpm of 329-339 (High Warning).

(1) Bright-Dim Switch The BRIGHT-DIM switch

(2) For rotor rpm of 300-310 (Low Warning).

on the CAUTION panel permits the pilot to manually

select a bright or dimmed condition for all the individual

(3) For engine rpm of

6100-6300 (Low

worded segments and the master caution indicator. The

Warning).

dimming switch position will work only when the pilot

instrument lights are on. The master caution system

(4) Loss of signal (circuit failure) from either

lights will be in bright illumination after each initial

rotor tachometer generator or power turbine tachometer

application of electrical power, when the pilot instrument

generator.

lights are turned OFF, or a loss of power from the dc

b. Warning light and audio warning signal

essential bus occurs.

combination:

(2) Reset-Test Switch The RESET-TEST

(1) For rotor rpm of 300-310 and engine rpm of

switch on the CAUTION panel enables the pilot to

6100-6300 (Low Warning).

manually reset and test the master caution system.

(2) Loss of signal (circuit failure) from both

Momentarily placing the switch in the RESET position,

rotor tachometer generator and power turbine

extinguishes and resets the master caution indicator light

tachometer generator.

so it will again illuminate should another fault condition

occur. Momentarily placing the switch in TEST position

will

2-18

Change 17

TM 55-1520-210-10

Rotor Tachometer Generator and Power

Powerplant ODDS Components.

Turbine Tachometer Generator. The rotor tachometer

generator and power turbine tachometer generator both

(1) Oil separator (Lubriclone) located in engine

send signals to the high-low rpm warning light and audio

service compartment.

warning circuits.

When the warning light only is

energized determine the cause of indication by

(2) Oil Filter, equipped with

3-micron filter

checking the torquemeter and cross referencing other

element, located in engine service compartment.

engine instruments. A normal indication signifies that the

engine is functioning properly and that there is a

(3) Chip detector located at bottom of oil

tachometer generator failure or an open circuit to the

separator. Detector is wired to Engine Chip Detector

warning system rather than an actual engine failure.

Caution Capsule.

Electrical power for system operation is supplied by the

28 Vdc essential bus.

b. Drive System ODDS Components.

d. Light - High Low Limit RPM Warning. The high

low warning light (fig 24) is located on the instrument

(1) Full-Flow Debris Monitor in transmission

panel. This light illuminates to provide a visual warning of

sump replaces wafer disc filter.

low rotor rpm low engine rpm or high rotor rpm.

e. Switch - LOW RPM A UDIO/OFF. The LOW

(2) External Filter with

3-micron element

RPM AUDIO/OFF switch is on the engine control panel

located in Cargo Sling Compartment (hell hole). Filter

(fig 2-6). When in OFF position the switch prevents the

replaces existing 25 micron external filter.

audio warning signal from functioning during engine

starting. Current production helicopters use a spring-

(3) Three chip detectors, one in debris monitor,

loaded switch. When the switch has been manually

and one each in 42 and 90 degree gearboxes. Detectors

turned off for engine starting It will automatically return to

are wired to the Chip Detector Caution Capsule.

the AUDIO position when normal operating range is

reached.

c. Electrical System Component. Power module

2-85.1. Oil Debris Detection System (ODDS).

located in overhead in cabin provides electrical power to

ODDS improves oil filtration and reduces nuisance chip

pulse (burn) away ferrous (iron, steel) debris less than

indications caused by normal wear particles on detector

0.005 inch in cross section. Larger debris will not pulse

gaps. When a chip gap is bridged by conductive

away, but bridges chip gap and closes the circuit to

particles, a power module provides an electrical pulse

caution capsule.

which burns away normal wear particles.

Section XIV. SERVICUNG PARKING AND MOORING

2-86. Servicing

a. Servicing Diagram Refer to figure 2-10.

c The helicopter is serviced as follows:

b. Approved Military Fuels, Oils and Fluids. Refer

(1) Refer to figure 2-10 for fuel filler location.

to table 2 -1.

(2) Assure that fire guard is in position with fire

c. Fuel Sample. Settling time for AVGAS is 15

extinguisher

minutes per foot of tank depth and one hour per foot

(3) Ground servicing unit to ground stake.

depth for Jet (JP) fuels. Allow the fuel to settle for the

(4) Ground servicing unit to helicopter.

prescribed period before any fuel samples are taken.

(5) Ground fuel nozzle to ground receptacle

Tank depth is about 29 inches.

located adjacent to fuel receptacle on helicopter.

2-87. Fuel System Servicing

Ensure that servicing unit pressure is not above 125

psi while refueling.

(6) Closed circuit.

Servicing personnel shall comply with all

(a) Remove fuel filler cap and assure that

safety precautions and procedures

refueling module is in locked position. Refer to figure 2-

specified in FM

1068 Aircraft Refueling

10.

Field Manual.

a. Refer to table 2-1 for fuel tank capacities.

b. Refer to table 2-1 for approved fuel.

Change 17

2-19

TM 55-1520-210-10

(b) Remove nozzle cap and insert nozzle into

2-88. Approved Commercial Fuel, Oils, and Fluids.

fuel receptacle and lock into position.

a. Fuels. Refer to table 2-1.

b. Oils. Refer to table 2-1.

position. Fuel flow will automatically shut off when fuel

c Fluids. Refer to table 2-1

cell is full. Just prior to normal shut off fuel flow may

2-89.

Use of Fuels.

cycle several times as maximum fuel level is reached.

a There are no special limitations on the use of

(d) Assure that flow control handle is in OFF or

Army standard or alternate fuels but certain limitations

NO FLOW position and remove nozzle.

are imposed when emergency fuels used. A fuel

(7) Gravity or open port:

mixture which contains over 10 percent leaded gasoline

(a) Remove fuel filler cap.

shall be recorded as all leaded gasoline. The use of

(b) Using latch tool attached to filler cap cable

emergency fuels shall be recorded m the

open refueling module

FAULT/REMARKS column of DA Form 2408- 13, Aircraft

Maintenance and Inspection Record noting the type of

fuel receptacle.

fuel, additives, and duration of operation.

(d) Fill for specified level.

b. When mixing of fuel in helicopter tanks or

(e) Remove nozzle

changing from one type of authorized fuel to another, for

(f) Close refueling module by pulling cable until

example, JP-4 to JP-5, it is not necessary to drain the

latch Is in locked position. Refer to figure 2-10.

helicopter fuel system before adding the new fuel

(8) Replace fuel nozzle cap.

Table 2-1

(9) Replace fuel filler cap.

Servicing Table of Approved Fuels, Oils, and Fluids.

System

Specification

(10)Disconnect fuel nozzle ground.

Fuel MIL-

T-5624 (JP4)a

(11)Disconnect ground from helicopter to servicing

Crashworthy System-

unit

Total: 208 5 U.S Gallons (789 2 liter).

(12)Disconnect servicing unit ground from ground

Usable: 206.5 U S gallons (781 6 liters)

stake.

Internal Auxiliary Tanks-

Usable 300 U S gallons (1135.5 liters)

(13)Return fire extinguisher to designated location.

Oil:

d. Rapid (Hot) Refueling.

Engine

MIL-L-2369934

(1) Before rapid refueling.

*MIL-L-780824

(a) Throttle - idle.

Transmission

MIL-L-2369934

(b) FORCE TRIM

ON or controls

DOD-L-8573434

*MIL-L-7808 24

frictioned.

42’ Gearbox

MIL-L-2369934

DOD-L-8513424

(2) During rapid refueling. A crewmember should

*MIL-L-780824

observe the refueling operation (performed by authorized

90’ Gearbox

MIL-L-2369934

DOD-L-8573424

refueling personnel) and stand fire guard as required.

*MIL-L-780824

One crewmember shall remain in the helicopter to

Hydraulic System

MIL-H- 560667

monitor controls. Only emergency radio transmission

*MIL-H-8328267

should be made during RAPID refueling.

Main Rotor Grip

MILL-4615267

MIL-L-2369934

(3) After rapid refueling, the pilot shall be advised

*MIL-L,780824

by the refueling crew that fuel cap is secure and

MIL,L-210489

grounding cables have been removed.

MIL-L-461789

Pillow Block Oil

MIL-L2369934

MIL-L7808 248

*MIL-L,-210489

MIL-L-4615289

MIL-L-4616789

FOOTNOTES

Army Standard fuel is MIL-T-5624 (IP-4) NATO code is F-40

Alternate fuels are MILT-5624 (IP-5) (NATO F-44) and MIL-T-83133

(P-8) (NATO F-34). Emergency fuel is MIL-G-5572 (any AV gas)

(NATO F-12, F-18, F-22). Refer to TM 55-9150-200-24.

The helicopter shall not be flown when emergency fuel has been used

for a total cumulative time of 50 hours The engine manual also limits

operation to 25 hours when TCP s m the fuel.

2-20

Change 17

TM 55-1520-210-10

It is not advisable to mix MIL-H-5606 and MIL-

83282 fluids, except during an emergency. An entry on

DA Form 2408-13 is required when the fluids are mixed.

* Lubrication oil made to MIL-L-7808 by

When changing

from MIL-H-5606 to MILH-83282,

Shell Oil Company under their part number

not more than two percent of MIL-H-5606 may be

307, qualification number 7D-1 shall not be

present in the system.

used in the engine or aircraft systems. It

Refer to stencil on grip assembly to determine

contains additives which are harmful to

proper lubrication requirements.

seals in the system.

MIL-2104, MIL-L-46152, and MILL-46167, must

be used in hub P/N 204-012-101-31 as follows:

2*MIL-L-7808 NATO code is 0-148. For use in ambient

Average Temp Range

Specifications

temperatures below minus 32’C/25’F. May be I used

+5°C and above

MILL-2104, Grade 40,

when MIL-L-23699 or DOD-L-85734 oil is not available.

NATO Code 0-230

Not for use m main rotor hub P/N 204-012-101-31

-18°C to + 5°C

MIL2104, Grade 30,

NATO Code 0-230 or

MIL-L6152, Grade 30

-29°C to -18°C

MILL-2104, Grade 10,

NATO Code 0-230 or

Under no circumstances shall MIL-L-23699

MIL-L46152, Grade

or DOD-L-85734 oil be used in ambient

10W30

temperatures below minus 32°C/25’F.

-54’C to -20°C

MILL-46167, DEXRON II

Aut6matlc transmission

MIL-L-23699 NATO code Is O-156 For use in

fluid.

ambient temperatures above minus 32’ C/25’F. Not for

Approved domestic commercial fuels (spec. ASTM- D-

use in main rotor hub P/N 204-012-101-31.

1655-70: Manufacturer’s designation-

Do not mix MI-LL-2104, MI-L-46152, MI L-

46167, MIL-L-23699, DOD-L-85734, and or MILL-

Jet B-JP4 Type

Jet A-JP5 Type

Jet A-1-JP8 Type

7808 oils, except during an emergency. If the oils are

mixed the system shall be flushed within six hours and

American JP4

American Type A

filled with the proper oil. An entry on DA Form 2408-13

AeroJet B

AeroJet A

AeroJet A-i

is required when the oils are mixed.

Richfield A

MILH-5606 NATO code is H-515. For use in

B.P.A.T.G.

B.PA.T K

ambient temperatures below minus 35’C/30’F (Refer to

Caltex let B

Caltex Jet A-1

TB 55-1500-344-25)

CITGO A

For use In ambient temperatures above minus

Conoco IJP-4

Conoco Jet-50

Conoco Jet-60

35°C/30°F.

Gulf let B

Gulf let A Gulf

Jet A-

EXXON Turbo Fuel EXXON A

EXXON A-I

Mobil Jet B

Mobil Jet A

Mobil Jet A-I

Philjet JP-4

PhilJet A-50

Prolonged contact with hydraulic fluid or

Aeroshell JP4

Aeroshell 640

Aeroshell 650

its mist can irritate eyes and skin. After

SuperJet A

Superjet A-1

any prolonged contact with

skin,

Jet A Kerosine

Jet A 1Kerosime

immediately wash contacted area with

Chevron B

Chevron A-50

Chevron A-1

soap and water. If liquid contacts eyes,

Texaco Avjet B

Avjet A

AvJet A-I

flush immediately with clear water. If liquid

Union IP4

76 Turbine Fuel

is swallowed, do not induce vomiting; get

immediate medical attention. When fluid is

decomposed by heating, toxic gases are

released.

Change 17

2-21

TM 55-1520-210-10

Approved foreign commercial fuels:

Country

F40

F-44

Brayco 899/899-G/899-S

Castrol 205

Belgium

BA-PF-2B

Jet Engine Oil 5

Canada

3GP-22F

3-6P-24e

STO-21919/STO-21919A/STD-6530

Denmark

JP-4 MIL-T-5624

HATCOL 3211/3611

France

Air 3407A

Turbo Oil 2380 (WS-6000)/2395 (W-6459)2392/2393

Germany (West)

VTL-9130-006

UTL 9130-007/

Mobil Jet II RM-139A/Mobil Jet II RM-147A/Avrex S

UTL 9130-010

Turbo260/Avrex S

Greece

JP4 MIL-T-5624

Turbo 265

Italy

AA-M-C-1421

AMC-143

Royco 899 (C-915)/899SC/Stauffer Jet II

Netherlands

JP-4 MIL-T-5624 D. Eng Rd 2493

Aeroshell Turbine Oil 500

Norway

JP4 MIL-T-5624

Aeroshell Turbine Oil 550

Portugal

JP-4 MILT-5624

Chevron Jet Engine Oil 5

Turkey

YP4 MIL-T-5624

Stauffer 6924/Jet I

Untied Kingdom

D. Eng Rd 2454

E. Eng Rd 2498

SATO 7377n730,TL-8090

(Britain)

NOTE: Anti-icing and Biocidal Additive for Commercial

Approved domestic commercial fluids for MIL-H-5606:

Turbine Engine Fuel - The fuel system icing inhibitor

Manufacturers designation-

shall conform to MIL-I-27686 The additive provides anti-

"PO" 4226

icing protection and also functions as a biocide to kill

Brayco 757B

microbial growths in helicopter fuel systems. Icing

Brayco 756C

microbial conforming to MIL-I-27686 shall be added to

Brayco 7561D

commercial fuel, not containing an Icing inhibitor, during

Hyspin A

refueling

operations

regardless

ambient

Univis J41

temperatures.

Refueling

operations

shall

Aero HFB

accomplished m accordance with accepted commercial

Petrofluid 5606B

procedures. Commercial product ’PRIST’ conforms to

Petrofluid 4607

MIL-I-27686.

Royco 756C/D

Royco 782

Approved domestic commercial oils for Mil-L-7808:

XSL 7828

Manufacturers designation-PQ Turbine Oil

8365

PED 3565

ESSO/ENCO Turbo Oil 2389 RM- 184A/RM-201A151

PED 3337

TL-5874

Aero Hydroil 500

YT-283

Do not use Shell Oil Co, part No.

307,

FP-221

qualification No. 7D-1 oil (MIL-L-7808). It

Approved domestic commercial fluids for MIL-H-83282:

can be harmful to seals made of silicone.

Brayco Micronic 882

Hanover R-2

Approved domestic commercial oils for MIL-L-23699:

HF 832

Manufacturers designation-PQ Turbine Lubricant

XRM 230A

5247/6423/6700/7731/8878/9595

XRM 231A

2-22

Change 17

TM 55-1520-210-10

VHF navigation (Omni) antenna

Aft position light (White)

Synchronized elevator

42 degree gearbox

Anti-collision light

HF long wire antenna

FM homing antenna No.1

Electrical compartment access door

Loop (ADF) antenna (Removed W/MWO 1-1620-210-50-30)

Aft radio compartment access doors

Position light (White)

Cabin door

Position light (Red)

Position Light (Red)

FM communications antenna No.2

Copilot door

VHF/UHF antenna

Static port

Pitot tube

WSPA Windshield Wiper Deflector

WSPS Windshield Deflector

Radio compartment and fwd battery location assess door

WSPS Upper Cutter

Pilot door

29A AN/ASN-175 antenna

Position fights (Green upper and lower)

90 degree gearbox

Heater compartment access door

Position Light (NVG)

Oil cook fan access door

Aft position light (NVG)

FM communications antenna No. 1

KY - 100 Processor

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

Change 18

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ARMY MODEL UH-1H/V HELICOPTERS. OPERATOR’S MANUAL (1988) - page 1