TECHNICAL MANUAL MAINTENANCE MANUAL FOR ARMY CH-47D HELICOPTER (1992)

TM 55-1520-240-10

When the right pedal is displaced forward, the forward

not apparent to the pilot because AFCS control inputs

rotor disk tilts to the right and the aft rotor disk tilts to

do not move the cockpit controls. The pitch, roll, and

the left. The opposite action occurs when the left pedal

yaw axis all operate in fundamentally the same manner,

is displaced forward. An ILCA is installed to assist the

Should a hardover occur, the pilot can easily override

pilot in moving the pedals.

AFCS.

The pedals are adjusted individually fore and aft by

f. Pitch attitude stability, airspeed hold, and a posi-

pressing a lever mounted on the pedal support and

tive stick gradient from hover to Vmax are provided

moving the pedal to a new position before repositioning

through the DASH actuator. The DASH actuator ex-

the lever. Insure that both pedals are adjusted equally

tends or retracts to maintain airspeed for a given stick

(left and right pedals in same respective hole position)

position.

and pedal adjustment lockpins are engaged. A balance

2-84. Bank Angle Hold.

spring is installed to stop pedal creep. A viscous damper

is installed to reduce control sensitivity.

Bank angle trim without cyclic stick movement is pro-

vided through left or right positioning of the cyclic stick

2-83. Advanced Flight Control System (AFCS).

AFCS trim switch. Bank angle hold is disengaged

anytime a CENTERING DEVICE RELEASE switch is

a. The Advanced Flight Control System (AFCS) sta-

pressed, a cyclic stick is moved laterally, or the HDG

bilizes the helicopter about all axes and enhances

switch is ENGAGED. Bank angle hold cannot be

control response. It automatically maintains desired

reengaged until the roll rate is less than 1.5° per second.

airspeed, altitude, bank angle, and heading. An auto-

matic turn feature, coupled to the pilot or copilot HSI

2-85. Heading Hold.

(horizontal situation indicator) is also included in the

AFCS.

The directional gyro provides an input to each AFCS

which signals the yaw ILCA to maintain heading within

b. Built In Test Equipment (BITE) is installed in

5 degrees. Heading hold is disengaged if the swivel

each AFCS computer. This equipment is intended for

switch is set to STEER or UNLOCK, a CENTERING

ground troubleshooting purposes only. An interlock

DEVICE RELEASE switch is pressed, or the direc-

circuit through the engine condition control box pre-

tional pedals are moved. Also, heading hold will be

vents BITE use anytime either ECL is out of STOP.

disengaged at airspeed

above 40

knots anytime lateral

c. Power is supplied to the HDG ENGAGED, BARO

trim is used, the stick is moved laterally, or the HDG

ALT and RAD ALT ENGAGED lights from the DC

switch is ENGAGED. Heading hold will not resume

essential bus through the CAUTION PNL circuit

until yaw rate is less than 1.5° per second at an airspeed

breaker on the No. 1 PDP. The No. 1 AFCS receives AC

above 40 knots with a bank angle of

less than 1.5°.

and DC power from the No. 1 AC and DC buses

respectively through the AFCS NO. 1 circuit breakers

2-86. Airspeed Hold.

on the No. 1 PDP. The No. 2 AFCS receives AC and

The airspeed hold feature provides a constant airspeed

DC power from the No. 2 AC and DC buses respectively

and pitch attitude relative to cyclic stick position at

through the AFCS NO. 2 circuit breakers on the No. 2

airspeeds above 40 knots. Airspeed and pitch can be set

PDP.

with the AFCS trim switch on the cyclic stick or by

d. The AFCS consists of the following components:

displacing the cyclic stick until the desired airspeed is

achieved then pressing the CENTERING DEVICE

(1) A cockpit control panel.

RELEASE switch. Refer to Chapter 8 AFCS Off Flight

(2) Two AFCS computers in the avionics com-

Characteristics.

partment.

2-87. Altitude Hold.

(3) Three ILCA’s in the flight controls closet.

Two methods of altitude hold can be selected. They are

(4) A differential airspeed hold (DASH) actua-

radar altitude hold or barometric altitude hold.

tor in the flight controls closet.

a. Radar Altitude Hold.

Radar altitude hold will

(5) Two longitudinal cyclic trim (LCT) actuators

maintain a more precise altitude in hover or over water

are installed, one in the forward upper controls, the

flight than barometric altitude hold. Maximum altitude

other in the aft upper controls.

for the use of radar altitude hold is

1,500 feet AGL.

(6) Roll and yaw magnetic brakes, a longitudinal

An error signal, caused by radar altitude deviations, is

CCDA, and a thrust CCDA are all located in the flight

derived from the pilot radar altimeter receiver-

controls closet.

transmitter and is processed by the No. 1 AFCS com-

(7) Three control position transducers.

puter. The processed error signal is applied to the

collective CCDA which drives the THRUST CONT

e. Attitude changes sensed by the attitude gyros, a

levers in the direction necessary to null the error signal.

yaw rate gyro in each AFCS computer, and the direc-

tional gyro are processed by the AFCS computers and

b. Barometric Altiude Hold.

Barometric altitude hold

applied to the ILCA’s. The ILCA’s extend or retract

is used in forward flight over terrain. It uses error

and move the upper flight controls. This control input is

signals produced within the No. 1 AFCS computer.

2-35

TM 55-1520-240-10

These error signals are in response to static pressure

changes and are proportional to altitude changes. The

signal is processed by the AFCS computer and applied

to the collective CCDA which drives the THRUST

CONT levers in the direction necessary to null the error

signal.

2-88. Heading Select,

Heading select is engaged when the HDG switch on the

AFCS panel (fig. 2-28) is pressed and the ENGAGED

light illuminates. The heading bug on the selected HSI

is the referenced heading. Rotating the HDG knob of

the HSI to set the bug at a new referenced heading

produces an error signal which is processed by the

AFCS computers and applied to the roll ILCA. The roll

ILCA then moves to produce a standard rate turn up to

a maximum bank angle of 2°

until the selected heading

is captured. Heading select can only be selected at

airspeeds above 40 knots. Heading select is disengaged

anytime a CENTERING DEVICE RELEASE switch is

pressed, the HDG switch on the AFCS panel is disen-

gaged, or when the opposite CMD SEL switch on the

Figure 2-28. Advanced Flight Control System

HSI MODE SELECT panel is pressed. .

Panel

2-89. Longitudinal Cyclic Trim System.

Heading select is disengaged if either CENTERING

Longitudinal cyclic trim (LCT) control is part of AFCS.

DEVICE RELEASE switch is pressed.

LCT reduces fuselage nose down attitude as forward

(2) BARO ALT and RAD ALT Switches. The

airspeed is increased, thus reducing fuselage drag. The

BARO ALT and RAD ALT are used to select altitude

system also reduces rotor blade flapping which results in

hold mode. An interlock prevents both switches from

lower stresses on the rotor shafts. The LCT actuators

being engaged at the same time. When pressed, the

are installed under the swashplates. Signals are trans-

ENGAGED legend will illuminate. RAD ALT hold is

mitted to these actuators either automatically by AFCS

used below 1,500

feet AGL. BARO ALT hold is used in

or manually by CYCLIC TRIM switches on the AFCS

forward flight to maintain a constant cruise altitude or

panel. Landing gear proximity switches drive the actu-

may be used in HOGE.

ators to GND (ground) operating position on ground

contact.

b. SYSTEM SEL Switch.

The SYSTEM SEL switch

is a five position rotary switch labeled OFF, 1, BOTH, 2,

2-90. Controls and Indicators.

OFF. Normally, the switch is at BOTH. In this position,

both AFCS are operating at

one-half gain. If one system

2-91. AFCS Control Panel. The AFCS control panel

should fail, the good system is selected and that system

(fig. 2-28) is on the canted console. It consists of the

operates at

3/4 gain. At OFF, both systems are inoper-

heading and altitude select, SYSTEM SEL (select), and

ative except for CYCLIC TRIM.

CYCLIC TRIM sections.

c. CYCLIC TRIM Switches.

The AUTO and MAN-

a. Heading and Altitude Select Switches.

The legend

UAL switch selects the mode of cyclic trim operation.

on these switches will dim when the PLT INST rotary

The FWD and AFT switches are used to extend or

control switch is placed out of the OFF detent.

retract the appropriate cyclic trim actuator.

(1) HDG Switch. The HDG (heading) switch is

(1) AUTO and MANUAL switch. A two-position

used in conjunction with the CMD SEL switch on either

switch which is normally placed in AUTO.

HSI MODE SELECT panel and the heading bug on

either HSI to select coupled turns. The switch can be

(a) AUTO Mode. In this mode, No. 1 AFCS

used only when airspeed is

above 40 knots. When the

controls the forward actuator and the No. 2 AFCS

switch is pressed and either CMD SEL switch is pressed,

controls the aft actuator.

the helicopter will automatically turn to and capture the

heading bug on the selected HSI. In addition, the

(b) MANUAL Mode. In this mode, the actu-

ENGAGED legend will illuminate. The switch is disen-

ators can be controlled with separate FWD and AFT

gaged by pressing it again.

actuator control switches, using the airspeed indicator

and the CYCLIC TRIM indicators.

Heading intercept will be at a standard rate of

3° per

second up to a bank angle limit of

20° at 133 knots. The

(2) FWD and AFT switches. Three-position

helicopter must be trimmed before engaging the mode

switches that can be placed in the EXT (extend) or RET

and cyclic stick control inputs should be avoided except

(retract) position. These switches are spring-loaded to

for longitudinal AFCS trim inputs to adjust airspeed.

the center off position. If the cyclic trim actuators fail to

2-36

TM 55-1520-240-10

extend or retract as indicated on the CYCLIC TRIM

indicators, MANUAL mode can be selected.

2-92. Cyclic Trim Indicators.

If the longitudinal cyclic trim actuators fail

at the full retract position or are manually

selected to the full retract position, do not

exceed the airspeed limitations shown in fig.

5-6.

The FWD and AFT CYC (cyclic) TRIM indicators (fig.

2-29) are on the center instrument panel. The indicators

are labeled 60 RET, GND, 150 EXT. The indicators

display position of the forward and aft LCT actuators

relative to airspeed. During ground operations, the

pointer will be at GND to indicate activation of the

landing gear proximity switches.

2-93. AFCS OFF Caution. Two AFCS OFF caution

capsules are on the master caution panel (fig. 2-51).

They are labeled NO. 1 AFCS OFF and NO. 2 AFCS

OFF. These cautions will illuminate when the associ-

ated AFCS is manually shutoff or has failed or the

associated DASH is in a low rate condition. Refer to

Chapter 8 AFCS Off Flight Characteristics.

2-94. Command Select Switch. The CMD SEL

switch is on the pilot and copilot HSI MODE SELECT

panels (Chapter 3). The switches are used to select the

HSI which will provide the referenced heading when the

HDG switch is engaged. Only one CMD SEL switch

may be selected at a time. If the other CMD SEL switch

is selected during heading select operations, the HDG

switch on the AFCS panel will disengage and heading

Figure 2-29. Forward and Aft Cyclic Trim

select will be disabled until the HDG switch is again

Indicators

pressed. When selected, the SEL legend on the switch

illuminates.

2-37

TM 55-1520-240-10

SECTION VI HYDRAULIC SYSTEMS

2-95. Hydraulic Power Supply System.

2-98. Utility Hydraulic System.

The hydraulic power supply systemThe utility hydraulic sy

stem supplies hy

rate systems. They are the No. 1 flight control system, No.

wheel brakes, power steering actuator, s

2 flight control system, and a utility system. Each system

ing cams, ramp actuating cylinders, hyd

includes a variable delivery pump and a reservoir cooler. In

motor, actuator for the center cargo h

addition, each flight control system has a power control

winch control valve, two engine starters

module, and the utility system has a pressure control

start circuit. When the APU is running,

module. Each flight control system is connected to the

system is pressurized by an APU driven

utility system by a power transfer unit (PTU). All systems

APU is not running and the rotors are t

are serviced by a common fill module and are pressurized to

hydraulic system is pressurized by an

prevent pump cavitation.

driven pump.

2-96. Flight Control Systems.

The utility hydraulic system incorporate

The No. 1 and No. 2 flight control systems are identical.

module which isolates utility subsystems

They are parallel in operation, hydraulically separated, and

When a failure occurs in one utility hy

electrically integrated. The flight control systems operate at

the remaining subsystems continue to op

approxi

3,000

psi, which is

500

psi for

the BRK STEER and RAMP PWR switches in t

ILCA operation. They power four dual upper boost actua-

are set to OFF.

(3,000

psi) and four

1,500

psi). Each flight

control system powers

one piston The APU starting.

subsystem of the utilit

No. 1 flight control system is princludes three accumulators which accele

forward transmission. No.

2 syststart, maintain reservoir pressure throu

pump on the aft transmission. Theand control operationsf the APU motor

consist of pressure-operated valstarting subsystem also

includes a two-s

pressure-line and return-line filcharging.

the APU start accumulators. The

control module is in the forward is normally recharged by

the APU motor-

control module is in the aft pylon. The accumulators

ditional accumulato

APU is started An ad

dampen low frequency pressure surges and provide stored

system provides for limited brake operat

hydraulic power for peak loads.

utility hydraulic system failure. The s

The PTU in each system allows ground checkout of the

flight control systems with the rotors stopped. Each PTU

consists of a pump driven by a hydraulic motor which is

pressurized by the utility hydraulic system. The PTU’s are

controlled by the PWR XFER 1 and 2 switches on the HYD

panel in the overhead switch panel.

2-97. FLT CONTR Switch.

The FLT CONTR (flight

control) switch is located on the HYD panel in the overhead

switch panel (fig. 2-30). It is a three-position center locked

switch labeled 2 ON, BOTH, and 1 ON. This switch can be

used to turn off one of the flight control systems. provided

the other one is operating. Turning off one of the flight

control hydraulic systems disables the corresponding AFCS

and causes the remaining AFCS to make full corrections. In

addition the respective AFCS OFF and HYD FLT CONTR

caution capsules will illuminate. The FLT CONTR switch

shall be set to BOTH during all flight conditions.

A1 BOTH. both solenoid valves are deenergized open and

both flight control systems are pressurized. When the FLT

CONTR switch is set to 1 ON, the two-way solenoid valve

on No. 2 power control module is energized closed. This

causes No. 2 pressure-operated valve to close. depressuriz-

ing No. 2 system. When the FLT CONTR switch is moved

to 2 ON, the two-way solenoid valve on No. 2 power control

module is deenergized open, and No. 2 system is pressur-

ized. Simultaneously, No. 1 solenoid valve closes and No. 1

system is turned off.

Figure 2-30. Hydraulics Control Panel

2-38

Change

TM 55-1520-240-10

has an accumulator to keep the swsystem pressure for normal brake an

the BRK STEER switch is OFF.

OFF, the brake and steering isola

isolating the brake and steering s

Normal operating pressure range for the utility hydraulic

ity systems. ON is the

syste

500

3500

psi. During APU operamaining util

is incr

approximately

3350

psi for engineOFF is used when there has been a h

(See table 2-3 for flight controlbrake or steering system. Setting th

capacities and fig.

2-54 for accprevents loss of system

fluid. Thi

sures.)

utility subsystems to continue to

brake system contains an accumulato

2-99. PWR XFER Switches.

The two-positiosystem operation in a hydraulic fa

XFER (power transfer) 1 and 2 swisystem also has a smalleccumulator

UTIL (utility) hydraulic portion locks locked with the system isolat

(fig. 2-30). Each switch is labelBRK STEER switch and va

lve is sup

switch is ON,

28-volt DC opens the normally closed

HYDRAULICS BRK STEER circuit breake

solenoid valve in the corresponding PTU and opens a valve

PDP.

in the pressure control module. This allows utility hydraulic

system pressure to operate the hydraulic motor pump on that

2-101.1. RAMP EMER Control Switch.

PTU, pressurizing the flight control hydraulic system.

Consequently, the flight controls can be operated on the

ground for maintenance and checks without tWARNINGrotors

turning.

When both switches are ON, No. 1 andThe RAMP EMER control switch is intendedntrol

hydraulic systems will be pressurizedforemergencyhusehonlyduringsmokeandmotor

pumps of the PTU’s supply pressure fumeeliminationlprocedures. Inadvertentop-

operation. When the switches are OFFerationhofthescargoerampandcargoadoores

are closed and the flight controls cfromnthetcockpitmay result in injury to

the rotors are turning. Power for thpersonnel or damage to equipment.s supplied

by the No.

2 DC bus through the HYDRAULICS PWR

XFER circuit breaker on the No. 2 PDP.

The momentary, guarded, three-posi

2-100. RAMP PWR Switch.

(ramp emergency) control switch is

hydraulic portion of the HYD control

switch allows the pilot, in an emer

WARNING

or lower the ramp to a partially o

closed position. The switch is labe

When the RAMP PWR switch is at OFF, be

sure the RAMP CONTROL valve is not

(down), and is spring loaded to the

moved from STOP. Operating the valve from

The switch is active only when the

STOP to UP or DN may cause the ramp to

set to EMERG. For up operation, the

free fall.

while the momentary switch is held

will stop as soon as the switch i

operation, the switch has a minimu

The RAMP PWR switch is on the lower right side of the

which allows the pilot to lower the

UTIL hydraulic portion of the HYD control panel (fig.

momentarily moving the switch to D

2-30). The switch has three positions labeled ON, OFF, and

releasing il. The downward ramp mo

EMERG. At ON, the ramp isolation valve in the utility

seconds after the switch is selecte

system pressure control module is open, allowing system

second timer circuit). If the ramp

pressure for normal ramp operation. At OFF, the ramp

osed pos

isolation valve is closed, isolatitongue) are in the fully cl

remaining utility systems. Thisselection of the DN position will p

system fluid if the ramp system fthe cargo door to be fully retracte

power is supplied to the RAMP EMEcan be further lowered in

5 second

ramp and cargo door to be openedtarily reselecting the DN position

cockpit. Power to operate the RAMPdownward motion of the ramp may be

isolation valve is supplied by thby momentarily setting the

RAMP EME

HYDRAULICS UTIL SYS CONT circuitposition. Thehe

ramp can also be low

No. 1 PDP.

more than

5 seconds) by holding th

position until the desired ramp lev

2-101. BRK STEER Isolation Switch.

The BRK

STEER isolation switch is on the HYD control panel (fig.

At the UP or DN positon, 28-volt DC

2-30). It is a guarded two-position switch labeled ON and

tive up or down solenoid on th

OFF. At ON, the brake and steering isolation valve in the

utility system pressure control module is open, allowing

Change

2-39

TM 55-1520-240-10

valve. The ramp control valve hanhand pump may be used to

operate the ram

position, and the ramp repositionNORMAL position of the,check valve is

electrical power is removed fromsystem is pressurized

by the APU or by t

solenoids. The ramp control valvpump. When the engines

and the APU are n

STOP position and the ramp remaicontrollable check valve is set to OPEN

Power for the switch is supplied handle is set to UP or DN, and the hand

bus through the RAMP PWR switch aWhenerampPmovement

is completed, the

CONT circuit breaker on the No. 1handl

e is set to STOP and the controllab

to NORMAL. This valve may also be used

2-102. Hydraulic System Service Module.

A ser-

event of utility pump or system failure

vice module, on the right

side of the cargo compartment

lator pressure

to the subsystems.

above the ramp, provides for filling the two flight control

hydraulic systems and the utility hydraulic system. It

105. Hydraulic Pressure Cautions.

Three hydrau-

consists of a filler assembly, a two-stage hand pump, and a

lic pressure caution capsules, one for

selector valve for selection of any of the three hydraulic

system and one for the utility hydraulic

systems for tilling.

on the master caution panel (fig.

2-51

2-103. Utility System Hand Pump.

A two-stage NO.

1 HYD FLT CONTR, NO. 2 HYD FLT CONT

pump, on the right side of the caUTIL HYD SYS. Each capsule is electrica

ramp, is used to pressurize the AaUpressurecswitch in the corresponding

APU starting. Also, in conjunctioWhenever hydraulicbelow

1,800supsi in one

PRESS controllable check valve, iofmathe flight control systems or the

the ramp and door.

system caution

illuminates. The cautio

guishes as incr

approaches

2,300

psi. Cau-

2-104. EMERG UTIL PRESS Controllable Check

tion capsule operation is independent o

Valve.

The EMERG UTIL PRESS controllable check

indicator operation. Power for these ca

valve is located above the hand pump. It allows APU start

the DC essential bus through the CAUTI

accumulator pressure to be used for operation of the ramp or

breaker on NO. 1 PDP.

any other subsystem (brakes, swivel locks, etc.). When the

APU motor pump or utility pump is2-106.tHydraulicrPressurenIndicators. is not

Three HY-

necessary to use the hand pump unless the accumulator is

DRAULICS PRESSURE indicators (fig. 2-34)

discharged.

hydraulic system, are on the MAINTENAN

When the accumulator is dischargRefer to Section IX Utility Systems.

PRESS controllable check valve in conjunction with the

2-40

Change

TM 55-1520-240-10

SECTION Vll POWER TRAIN SYSTEM

2-107. General.

from the sump through the main lube pump, main filter,

cooler, and the jet protection screen to jets where the oil

Engine power is supplied to the rotors through a

is sprayed onto the various gears and bearings. In

mechanical transmission system (fig. 2-1). This system

addition, after the oil leaves the jet protection screen,

consists of a forward, a combining (mix), an aft, two

alternate paths routes some of the lubricating oil to the

engine transmissions, and drive shafting. An overrun-

aft shaft bearing and cooling oil to the generators.

ning sprag clutch is installed in each engine transmis-

Auxiliary system oil flows from the auxiliary sump

sion. The clutch provides a positive drive connection to

through the auxiliary pump and filter to the various

transmit power and permits freewheeling of both rotors

gears and bearings. Separate oil jets are utilized for

when in an actual autorotation or during a simulated

each oil system. The auxiliary system does not lubricate

power failure. Because of the freewheeling feature, no

the aft shaft bearing or the generators. An oil cooler

drag will be place on the rotors if an engine (or engines)

mounted on the aft end of the transmission cools main

fails.

system oil. Cooling air is drawn through the cooler by a

transmission-driven fan.

Power from the engine transmissions is transmitted

through separate drive shafts to the combining (mix)

2-111. Combining and Engine Transmission Lubri-

transmission. The combining (mix) transmission com-

cation Systems. The combining (mix) transmission

bines the power of the engines and transmits it at

contains the oil reservoirs to supply lubrication oil to the

reduced shaft speed to the forward and aft transmis-

various gears and bearings in the combining (mix)

sions. Further speed reduction occurs within the rotor

transmission, No. 1 engine transmission, and the No.

transmission.

engine transmission. Two lubricating pumps with four

elements each are within the combining (mix) transmis-

Two AC generators, the No. 2 flight control hydraulic

sion: left pump assembly and right pump assembly. The

pump, and the utility system pump are mounted on and

left pump assembly provides main lubrication to the

driven by the aft transmission. The No. 1 flight control

combining (mix) transmission and the No. 1 engine

hydraulic pump is mounted on and driven by the

transmission. The right pump assembly provides auxil-

forward transmission.

iary lubrication to the combining (mix) transmission and

2-108. Transmission Lubrication Systems.

lubricates the No. 2 engine transmission. Each pump

assembly contains two pumping elements and two scav-

The forward, aft, and combining (mix) transmissions

enge elements.

have independent main and auxiliary lubrication sys-

tems which operate concurrently. Each transmission has

Combining transmission main lubrication oil flows from

a filter with an impending bypass indicator. If the

the combining (mix) transmission oil reservoir through

differential pressure across the filter exceeds

15 to 18

the left pump assembly, filter, cooler, jet protection

psi, the bypass indicator will extend to indicate a

screen, and to the jets which spray the oil onto the

partially clogged filter. When the differential pressure

various gears and bearings. One of the scavenge ele-

reaches

25 to 30 psi, lubrication oil will bypass the filter.

ments of the left pump assembly returns the oil from the

Refer to table 2-3 for transmission oil system capacities,

combining (mix) transmission sump to the combining

oil specifications, and servicing procedures.

(mix) transmission oil reservoir. The auxiliary lubrica-

tion oil flows from the combining (mix) transmission

2-109. Forward Transmission. The forward trans-

auxiliary oil reservoir to the right pump assembly,

mission lubrication system supplies lubricating oil to the

auxiliary lubrication filter, and to the jets which spray

gears and bearings in the forward transmission. Main

the oil on to the various gears and bearings. One of the

system oil flows from the sump, through the main oil

scavenge elements of the right pump assembly returns

pump, oil filter, cooler, and a jet protection screen to

the oil from the combining (mix) transmission sump to

jets from which the oil is discharged to the various gears

the combining (mix) transmission oil reservoir. The

and bearings. Auxiliary system oil flows from the auxil-

right pump assembly does not route oil through a

iary sump through the auxiliary oil pump, and the

cooler.

auxiliary system filter to separate auxiliary oil jets. An

No. 1 engine transmission oil flows from the No.

oil cooler mounted on the aft end of the transmission

engine transmission oil reservoir on the combining (mix)

around the input pinion cools main system oil. Air is

transmission through the left pump assembly, filter,

forced through the cooler by a transmission-driven fan.

cooler, jet protection screen, and to the jets which spray

2-110. Aft Transmission. The aft transmission lubri-

the oil onto the various gears and bearings. One of the

cation system supplies lubricating oil to the various

scavenge elements of the left pump assembly returns the

gears and bearings in the aft transmission and to the aft

oil from the No. 1 engine transmission sump through a

rotor shaft bearing. In addition, the main lubrication

debris indicating screen and back to the No. 1 engine

system circulates cooling oil through the two AC gen-

transmission oil reservoir. No. 2 engine transmission oil

erators on the aft transmission. Transmission oil flows

flows from the No. 2 engine transmission oil reservoir

2-41

TM 55-1520-240-10

on the combining (mix) transmission through the right

pump assembly, filter, cooler, jet protection screen, and

to the jets which spray the oil onto the various gears and

bearings. One of the scavenge elements of the right

pump assembly returns the oil from the No. 2 engine

transmission sump through a debris indicating screen

and back to the No. 2 engine transmission oil reservoir.

Engine transmissions do not have auxiliary lubrication

systems.

All No. 1 and No. 2 engine transmission lubrication

system components are on the combining (mix) trans-

mission except the jet protection screens and jets.

Separate oil jets are utilized for each transmission

lubrication oil system. The individual oil coolers for the

combining (mix) and both engine transmissions are

mounted on the combining (mix) transmission and

utilize a common transmission driven fan for cooling air.

2-112. Transmission Main Oil Pressure Indicator.

A transmission main oil pressure indicator is located on

the center instrument panel (fig. 2-31). It indicates

either the lowest main oil pressure in any one of the

transmissions or only the oil pressure in the transmis-

sion selected by the pilot. The indicator is electrically

connected to each transmission. In addition, each trans-

mission and the aft rotor shaft bearing has a separate

low pressure switch. These switches are connected to

the XMSN OIL PRESS caution capsule on the master

caution panel and the TRANSMISSION MAIN OIL

PRESS indicating lights on the MAINTENANCE

PANEL (fig. 2-34). Power to operate the indicator is

supplied by the No. 1 AC bus through the XMSN OIL

PRESS circuit breaker on the No. 1 PDP.

2-113. Transmission Main Oil Pressure Selector

Switch. A transmission oil pressure selector switch is

located on the center instrument panel (fig. 2-31). The

switch positions are labeled TEST, SCAN, FWD, AFT,

MIX, LEFT, and RT. When the switch is set to TEST,

Figure 2-31. Transmission Main Oil Pressure In-

the pointer on the transmission pressure indicator will

drop to zero or below. When the switch is set to SCAN,

dicator and Selector Switch

the lowest main oil pressure among all the transmission

Each temperature probe incorporates a high oil temper-

will be indicated. The remaining positions are used to

ature switch which is independent of the temperature

select a particular transmission oil pressure indication.

indicator and is triggered at

140°C, lighting the XMSN

When selecting a particular switch position, be sure the

OIL HOT caution capsule on the master caution panel

switch is in detent. If the switch is not in detent, the

and to the TRANSMISSION OVERTEMP magnetic

pressure gage will indicate zero.

indicators on the MAINTENANCE PANEL. Power to

operate the indicator is supplied by the No. 1 AC bus

2-114. Transmission Main Oil Temperature Indica-

through the XMSN OIL TEMP circuit breaker on the

tor. A transmission oil temperature indicator is lo-

No. 1 PDP.

cated on the center instrument panel (fig. 2-32). It reads

from

-70° to + 150°C.

It indicates the highest oil

2-115. Transmission Main Oil Temperature Selec-

temperature among all the transmissions or only the oil

tor Switch. A transmission oil temperature selector

temperature of the selected transmission. A tempera-

switch is on the center instrument panel below the

ture probe is located in the forward and aft transmission

transmission oil temperature indicator (fig. 2-32). The

sumps and in each compartment of a three-compartment

switch positions are labeled TEST, SCAN, FWD, AFT,

oil tank for the combining (mix) transmission and in

MIX, LEFT, and RT. When the switch is set to TEST,

each engine transmission. The temperature probes in

the pointer on the transmission oil temperature indica-

the three tank compartments measure oil temperature

tor deflects full scale toward low temperature. When

in the tank and may not immediately indicate a trans-

the switch is set to SCAN, the highest oil temperature

mission problem. Loss of oil or low oil pressure may not

among all transmissions is indicated. The remaining

be accompanied by a high oil temperature indication.

positions are used for selecting a particular transmission

2-42

TM 55-1520-240-10

and indicator and the TRANSMISSION OVERTEMP

magnetic indicators on the MAINTENANCE PANEL.

b. XMSN OIL PRESS Caution.

It illuminates when

main oil pressure drops

below 20

psi in any transmission

or aft rotor shaft pressure drops

below

psi. The

low-pressure system is identified by the transmission oil

pressure selector switch and indicator on the center

instrument panel and the TRANSMISSION MAIN

OIL PRESS indicating lights on the MAINTENANCE

PANEL. If the XMSN OIL PRESS caution capsule

illuminates and the affected transmission cannot be

determined using the selector switch, the condition may

be caused by loss of aft rotor shaft oil pressure. Low oil

pressure at the aft rotor shaft is indicated by the

illumination of the TRANSMISSION AFT SHAFT

MAIN OIL PRESS indicating light on the MAINTE-

NANCE PANEL.

c. XMSN AUX OIL PRESS Caution.

It is activated

by individual aux oil switches and illuminates when

auxiliary oil pressure drops

below 20 psi in the fwd or aft

transmission and

10 psi in the combining (mix) trans-

mission. The transmission with the low pressure is

identified by a lit TRANSMISSION AUX OIL PRESS

indicating light on the MAINTENANCE PANEL.

d. NO. 1 or NO. 2 ENG XMSN HOT Caution.

They

illuminate if oil temperature in either engine transmis-

sion exceeds about

190°C. The capsules are activated by

a thermoswitch in each engine transmission. The ther-

moswitch monitors oil temperature in the transmission,

not in the reservoir. It is part of a chip detector and

temperature assembly in each engine transmission.

NOTE: REFER TO CHAPTER 5 FOR INSTRUMENT MARKING.

D145-82-10

2-117. Transmission Chip Detectors. Chip detec-

Figure 2-32. Transmission Main Oil Temperature

tors are installed in all transmission and aft rotor shaft

Switch and Indicator

thrust bearing lubrication systems. All transmission chip

detectors, except those in the engine transmission, are

oil temperature indication. When selecting a particular

connected to the XMSN CHIP DET caution capsule on

switch position, be sure the switch is in detent. If the

the master caution panel. Engine transmission chip

switch is not in detent, the oil temperature indicator will

detectors are connected to the corresponding NO. 1 or

indicate -70°C.

NO. 2 ENG CHIP DET caution capsules.

2-116. Transmission Oil Cautions. Five transmis-

All transmissions and the aft rotor shaft chip detectors

sion oil caution capsules are on the master caution

are also connected to the TRANSMISSION CHIP

panel. The capsules are labeled XMSN OIL HOT,

DETECTOR magnetic indicators on the MAINTE-

XMSN OIL PRESS, XMSN AUX OIL PRESS, and

NANCE PANEL. When a chip detector is bridged by

NO. 1 AND NO. 2 ENG XMSN HOT. These cautions,

ferrous particles, the XMSN CHIP DET, or the NO. 1

in conjunction with the transmission oil pressure and

and/or NO. 2 ENG CHIP DET caution capsule illumi-

nates. At the same time, the corresponding TRANS-

temperature indicators on the center instrument panel

MISSION CHIP DETECTOR indicator on the MAIN-

and the TRANSMISSION OVERTEMP magnetic in-

TENANCE PANEL will trip and change from an

dicators, MAIN OIL PRESS, and AUX OIL PRESS

all-black indication to a black-and-white indication,

indicating lights on the MAINTENANCE PANEL,

identifying the transmission.

alert the crew to impending transmission lubrication

problems. The cautions operate independently of the

2-118. Transmission Chip Detectors Fuzz Burn-

pressure and temperature indicators on the center

Off. Helicopters equipped with the chip detector fuzz

instrument panel.

bum-off system in the forward, combining (mix), aft,

No. 1 and No. 2 engine transmission, and aft rotor shaft

a. XMSN OIL HOT Caution.

It illuminates when

thrust bearing are identified by a module labeled PWR

the main oil temperature in the sump of the forward,

MDL CHIP BURN-OFF located below the MAINTE-

aft, and reservoir of the combining (mix) or either

NANCE PANEL. The chip detector fuzz burn-off

engine transmission

exceeds 140°C.

The hot transmis-

system employs an automatically operated fuzz bum-off

sion is identified by the oil temperature selector switch

electrical circuit with the ability to eliminate nuisance

2-43

TM 55-1520-240-10

SECTION Vlll

ROTOR SYSTEM

2-119. General.

of the shock absorber is disconnected, the blade can be

folded in either direction about the vertical hinge pin.

Lift is produced by a rotor system consisting of two fully

articulated counter-rotating rotors. Each rotor has three

2-120. Rotor Blades.

fiberglass blades. The forward rotor is driven by the

forward transmission through a rotor drive shaft. The

a. Each rotor blade consists of a D-shaped fiberglass

aft rotor is driven by the aft transmission through a

spar assembly and a Nomex fairing assembly bonded to

vertical drive shaft.

the spar. The blade chord is

32 inches.

The rotor head consists of a hub connected to three

b. A titanium nose cap is bonded to the leading edge

pitch-varying shafts by three horizontal hinge pins.

of the spar. A nickel erosion cap is bonded to the blade

These pins permit blade flapping. Stops on the top and

along the outer

54 inches of leading edge. This cap

the bottom of the hub limit the blade flapping motion.

protects the part of the blade most vulnerable to

The aft rotor head is equipped with centrifugal droop

erosion.

stops which provide increased blade flapping angle for

c. The fairing assembly is bonded to the trailing edge

ground and flight operation.

of the spar. These fairings are constructed of a Nomex

Covers may be installed on the centrifugal droop stop

honeycomb core covered with a fiberglass skin. Wire

operating mechanism. The covers prevent ice accumu-

mesh screens are embedded in the fiberglass skin at the

lation on the mechanism and ensure proper droop stop

tip and the trim tab. The wire mesh screens provide an

operation following flight in icing conditions. For infor-

electrical path to the rotor hub from the metal trim tab

mation on use of the droop stop covers, refer to Chapter

and tip for lightning protection. Also, to provide light-

8, Section IV.

ning protection, each blade has two lightning protection

cables and two straps. The cables and straps complete

Mounted coaxially over the pitch-varying shafts are

the path from the wire mesh to the rotor head.

pitch-varying housings to which the blades are attached

d. Balance and tracking weights are installed in the

by vertical hinge pins. These pins permit blade leading

and lagging. Each pitch-varying shaft is connected to the

tip of spar and fairing assembly. The tracking weights

pitch-varying housing by a laminated tie bar assembly.

are removable and are used for blade track and balance.

The high tensile strength and low torsional stiffness of

2-121. Rotor Tachometers.

the tie bar retains the blade against centrifugal force

and allows blade pitch changes about the pitch axis.

Two rotor tachometers (16, fig. 2-8 and fig. 2-10), one

mounted on the pilot instrument panel, the other

Blade pitch changes are accomplished by three pitch-

mounted on the copilot instrument, indicate percent of

varying links connected from the rotating ring of the

rotor revolutions per minute (RRPM). A small needle

swashplate to the pitch-varying housing on each rotor

on the tachometer indicates percent RPM from

0 to 60.

blade. Cyclic pitch changes are accomplished by tilting

The large needle indicates percent RPM from

60 to 130.

the swashplate. Collective pitch changes are accom-

The RRPM sense signal is supplied by the AC genera-

plished by vertical movement of the swashplate. Com-

tors. Generator No. 1 supplies the copilot indicator and

bined collective and cyclic pitch changes result from

generator No. 2 supplies the pilot indicator. Power to

combined control inputs by the pilot.

operate the indicators is supplied by the DC essential

A direct-action shock absorber is attached to the blade

bus through the ROTOR TACH circuit breaker on the

and to the pitch-varying housing. When the inboard end

No. 1 and No. 2 PDP.

2-45

TM 55-1520-240-10

SECTION IX UTILITY SYSTEMS

2-122. Anti-Icing Systems.

CAUTION

Anti-icing is provided for the pitot tubes, AFCS yaw

If windshield bubbling or delamination oc-

ports, and windshields.

curs around the sensor element, immediately

place switch to OFF for that windshield.

2-123. ANTI ICE Panel.

The ANTI ICE panel is located on the overhead switch

When any switch is moved to ON, current flows to the

associated temperature controller and then to the wind-

panel (fig. 2-33). It has three two-position W/S (wind-

shield. As the temperature of the windshield rises to a

shield) switches labeled CPLT, CTR, and PLT. The

preset value (about 44°C),

as sensed by the sensor

switches positions are OFF and ON. In addition, a

element, the electrical current to the windshield is

two-position PITOT heat switch is in this panel. The

interrupted by the temperature control relay. Once the

switch positions are OFF and ON.

windshield has cooled sufficiently, electrical current is

Power for the pilot and center windshields is from the

reapplied. This causes a cycling effect which maintains

No. 2 AC bus through the WSHLD ANTI ICE HEAT

windshield temperature within operating limits.

PILOT and CTR circuit breakers. Power for the copilot

Operating temperature is reached in

less than 1 minute

windshield is from the No. 1 AC bus through the

after the switch is placed to ON. When the switch is

WSHLD COPLT HEAT circuit breaker on the No. 1

placed to OFF, the anti-icing system is deenergized.

PDP. Anti-ice control for the pilot and center wind-

b. PITOT Heal Switch.

Heating elements prevent ice

shield is from the 28-volt No. 2 DC bus through the

accumulation in the pitot tubes and the yaw ports.

WSHLD ANTI ICE CONT CTR and PILOT circuit

When the PITOT switch is placed to ON, power to the

breakers on the No. 2 PDP. Anti-ice control for the

heater elements in the pitot tubes and yaw ports is

copilot windshield is from the 28-volt No. 1 DC bus

applied. When the switch is placed to OFF, the heating

through the WSHLD COPLT CONT circuit breaker on

elements are deenergized.

the No. 1 PDP. Power to operate the heater elements in

the pitot tubes and yaw ports is supplied by the No. 2

2-124. MAINTENANCE PANEL.

AC bus through the PITOT HEAT and YAW PORT

The MAINTENANCE PANEL is on the right side of

HEAT circuit breakers on the No. 2 PDP.

the cabin above the ramp (fig. 2-34). The panel is

provided to assist in the identification of system mal-

a. W/S Switches.

The pilot and copilot windshields

are anti-iced and defogged electrically. The center

function or condition that may require servicing or other

maintenance. The panel is divided into four sections.

windshield is defogged but not anti-iced. The laminated

They are labeled TRANSMISSION, HYDRAULICS,

windshield panels are heated electrically by current

ENGINE, and GROUND CONTACT.

which passes through a transparent conductive coating

embedded between the layers.

2-125. TRANSMISSION Section. This section mon-

itors the FWD, COMB, AFT, AFT SHAFT, LEFT, and

RIGHT transmissions. It consists of six CHIP DETEC-

TOR magnetic indicators, six DEBRIS SCREEN mag

netic indicators, six MAIN OIL PRESS indicating

PRESS-TO-TEST lights, three AUX OIL PRESS indi-

cating PRESS-TO-TEST lights, and five OVERTEMP

magnetic indicators. Power to operate the indicators is

supplied by the No. 1 DC bus through the HYDRAU-

LICS MAINT PNL circuit breaker on the No. 1 PDP.

a. CHIP DETECTOR Magnetic Indicators.

When the

corresponding CHIP DETECTOR is bridged by ferrous

particles, the associated chip detector indicator changes

from all-black to black-and-white. In addition, the

XMSN CHIP DET or ENG CHIP DET caution capsule

illuminates on the master caution panel.

b. DEBRIS SCREEN Magnetic Indicators.

There is

one indicator each for the FWD transmission, AFT

transmission, and LEFT and RIGHT engine transmis-

sions. There are two indicators for the COMB transmis-

sion. One indicator for the left sump and one indicator

Figure 2-33. Anti Ice Panel

for the right sump.

2-46

TM 55-1520-240-10

NOTE:

REFER TO CHAPTER 5 FOR INSTRUMENT LIMIT MARKINGS.

12501

Figure 2-34. Maintenance Panel

NOTE

XMSN OIL HOT caution illuminates on the master

Their is no cockpit indication of a latched

caution panel and trips the corresponding OVERTEMP

DEBRIS SCREEN magetic indicator. If a

magnetic indicator on the MAINTENANCE PANEL,

DEBRIS SCREEN magnetic indicator

thus identifying the hot transmission.

latches, the flight engineer shall advise the

2-126. HYDRAULICS Section. This section monitors

pilot immediately.

the FLT CONT NO 1, FLT CONT NO 2, and UTIL-

ITY hydraulic systems. It consists of three PRESSURE

The indicators are electrically connected to screens in

indicators, three fluid TEMPERATURE indicators,

the sumps of each transmission. If the screen mesh is

two RESERVOIR LEVEL indicators, six FILTER

bridged with conductive particles, the indicating circuit

CHANGE indicating PRESS-TO-TEST lights, and four

closes and trips the corresponding DEBRIS SCREEN

PUMP FAULT indicating PRESS-TO-TEST lights.

magnetic indicator on the MAINTENANCE PANEL.

Power to operate the indicators is supplied by the No. 2

DC bus through the HYDRAULICS MAINT PNL LTS

c. MAIN OIL PRESS Indicating Lights.

If main oil

circuit breaker on the No, 2 PDP.

pressure drops

below 20 psi in any transmission or 10 psi

in the aft shaft bearing, the corresponding indicating

a. PRESSURE Indicators.

The FLT CONT NO

light will illuminate. In addition, the XMSN OIL PRESS

and NO 2 PRESSURE indicators are electrically con-

caution will illuminate on the master caution panel.

nected to a corresponding pressure transmitter on the

respective power control module. The UTILITY PRES-

d. AUX OIL PRESS Indicating Lights.

If auxiliary oil

SURE indicator is electrically connected to a pressure

pressure drops

below

psi in the FWD or AFT

transmission or

10 psi in the COMB transmission, the

transmitter on the pressure control module. Indicator

corresponding indicating light will illuminate. In addi-

operation is independent of caution capsule operation.

tion, the XMSN AUX OIL PRESS caution will illumi-

Power to operate the indicators is supplied by the No. 2

nate on the master caution panel.

DC bus through the HYDRAULICS PRESS IND

circuit breaker on the No. 2 PDP.

e. OVERTEMP Magnetic Indicators.

Each OVER-

TEMP magnetic indicator is electrically connected to a

b. TEMPERATURE Indicators.

The indicators are

temperature probe in the reservoir of each transmission.

below the PRESSURE indicators. They indicate the

If oil temperature in the transmission reservoir,

exceeds

temperature of the hydraulic fluid at the outlet of the

140°C, a switch closes. When the switch closes, the

corresponding reservoir-cooler. Power to operate the

2-47

55-1520-240-10

2 DC bus throThis section consists of two indicating l

HYDRAULICS FLUID

TEMP circuit breaWhen the landing

gear proximity switch

PDP.

appropriate GROUND CONTACT indicating l

illuminate.

c. RESERVOIR LEVEL

The left indicator

is dedicated to the No. 1 and No. 22-129.GNDSwitch.ntrol hyraulics sys-

tem. In addition. a two-position FLT CONT switch labeled

NOTE

NO I and NO 2 is used to select the system of which the fluid

level is to be indicated. The reservoiWhile in flight. the flight

engineer s

FULL mark before flight. The right inpilot when placing theo

GND switch on t

the utility hydraulic system. When tTENANCE PANEL to TEST. Placing the swi

CHECK switch is pressed, the fluid lto TEST will cause the

NO. 1 and NO.

cooler will be indicated by the approCHIP DET, XMSN O

IL HOT, and XMSN CHIP

DET cautions to illuminate.

d. FILTER CHANGE indicating Lights. The indicat-

ing lights are arranged in three sThe GND switch allows the flight engineer

system. Each set of indicating lig(Built In Test Equipment)

test on the cir

RTN. The PRESS indicating light iTENANCE PANEL. The switch is springloaded

pressure line filter in each systeat center-off position. AttTEST, a black

monitors the return line filter appears on all magnetic BITE indicators.

pressure drop a

exceeds

psi. indicneti

ISsupplied

impending filter bypass, the corto the switch by the No.

1 DC bus through

indicating light will illuminate.MAINT PNL circuit breaker No. 1 PDP.

change indicating lights is supplied by the No.

2 DC bus

2-130. Windshield Wipers.

through the HYDRAULICS MAINT PNL LTS circuit

breaker on the No. 2 PDP.

CAUTION

e. PUMP FAULT IndicaTheg indicating

To prevent windshield damage, do not operate

lights are labeled NO.

1, NO.

2, APUwindshield wipers when windshields are dry.e

connected to sensors in the case drain line of each pump. If

the flow rate from the case drainTwo electrically driven windshield wiper

point which causes an increased pinstalled, one on eache

pilot windshield.

sensor, the sensor turns on the coboth wipers through two flexible shafts

light (a high flow rate from the wipers converters. The windshield wiper m

indicate impending pump failure). by the W/S (windshield) WIPER switch loca

is supplied by the No. 1 DC bus thhead switch panel.CS

Power is supplied by

MAINT PNL circuit breaker on the NthroughP.

the WSHLD WIPER circuit breaker

PDP.

2-127. ENGINE CHIP DETECTOR Section.

This

five positions

section consists of two magnetic iThe W/S WIPER switch hasd

and PARK. Wiper speed can

NO. 2. When the corresponding ENGISLOW, MED, FAST,

as desired, by rotating the switch from

is bridged by ferrous particles, the associated chip detector

ately at any posi

indicator changes from all-blackwipers will stop immedi

pers stop and rep

addition. the ENG CHIP DET cautiontravel. At PARK. the wi

inside windshield frame.

master caution panel.

2-131. Map and Data Case.

2-128. GROUND CONTACT Section.

The map and data case is in the passagewa

CAUTION

maps, and other data.

Should either or both GROUND CONTACT

2-132. Cockpit Rearview Mirror.

indicating lights remain illuminated after lift-

off to hover, the indicated system(s) DASH will

A rearview mirror is installed on the ri

not function properly in forward flight. If both

support to enable the pilot to observe th

GROUND CONTACT indicating lights

2-133. Spare Lamp Stowage Box.

remain illuminated after lift-off, the AUTO

function of both cyclic trims system will be

The spare lamp stowage box is in the cockpi

inoperative.

1 PDP. Spare lamps are provided for the ins

2-48

Change

TM 55-1520-240-10

instrument light shields, dome lig2-136.DC Cabin Utility Receptacles.

and nacelle work lights.

Four 28-volt DC utility receptacles

the sidewalls of the cargo compartm

2-134. Cockpit Utility Receptacles.

left cabin utility receptacles is s

Two 28-volt DC utility receptaclesthrough the UTILITY LH FWDnnd LH A

No. 1 PDP and one on No. 2 PDP. Eaon the No. 1 PDP. Power

to operate

UTIL RCPT 28V DC. Power to operatereceptacles is supplied by the No.

is supplied by the No. 1 DC bus thrUTILITY RCPT RH AFTTand RH FWD cir

circuit breaker on the No. 1 PDP.the No. 2 PDP.ate the pilot

receptacle is supplied by the No.

2 DC bus through the

2-137. Ash Trays.

UTILITY RCPT PILOT circuit breaker on the No. 2 PDP.

Three ash trays are installed in th

and one for the troop commander.

2-135. AC Cabin Utility Receptacles.

2-136. Compass Correction Card Holder.

A 115 volt, single-phase 400 Hz AC utility receptacle and a

The magnetic compass correction car

200 volt 3-phase 400 Hz AC utility receptacle are on each side

the left side of the magnetic compa

of the cabin at sta 320. The receptacles are accessible after the

necessary deviation values which ar

acoustical access cover and the receptacle dust cap are

reading.

removed. Power to operate the 115 volt receptacles is sup-

plied by the No. 1 and No. 2 AC buse2-139.P

ilot Assist Straps.

UTIL RCPT circuit breakers on No. Two assist straps are

attached to t

to operate the 200 volt receptaclesthe cockpit structure to provide t

No. 2 AC buses through the LH andwhile getting into the seats. The a

3-phase circuit breakers on No. 1

flat against the structure after us

Change

2-49

TM 55-1520-240-10

SECTION X HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL CONTROL

SYSTEMS

2-140. Heating and Ventilating System.

A 200,000 btu/hr capacity internal combustion heating

system is provided. It consists of a heater unit, a fuel

control unit, an ignition assembly, a blower, control

relays, and air pressure and temperature control cir-

cuits. Ducting carries heated air or ventilating air to the

cockpit and the cabin. The heater consumes approxi-

mately

15 pounds of fuel per hour from the right main

fuel tank.

The heater and blower are mounted vertically on the

right side of the helicopter, immediately aft of the

forward cabin section bulkhead. Air for the system is

provided by the blower which draws air from an inlet on

the forward upper side of the fuselage. If sufficient air is

not available for proper heater operation, an automatic

differential pressure switch in the heater circuit will stop

the heater.

Both ventilating and combustion air enters the heater

inlet. The heating air passes over the heated metal walls

of the combustion chamber and is directed to a network

of ducting. The air entering the combustion chamber is

combined with atomized fuel and, after combustion that

heats the metal walls, the exhaust is discharged through

an outlet on the forward upper side of the fuselage.

Power to operate the blower is supplied by the No. 2 AC

bus through the CABIN HEATER BLOWER circuit

breaker on the No. 2 PDP. Power to the rest of the

system is supplied by the No. 2 DC bus through the

CABIN HEATER CONT circuit breaker on the No. 2

PDP.

2-141. HTG Panel.

The HTG (heating) panel (fig. 2-35) is located on the

Figure 2-35. HTG Panel

overhead switch panel (fig. 2-12). It consist of a rheostat-

type CABIN TEMP SEL rotary switch, a three-position

The circuit to the thermostat heater winding is also

heater function switch, and a spring-loaded pushbutton

interrupted, allowing the winding to cool and the mer-

HTR START switch.

cury column to contract, thus reenergizing the temper-

ature controller relay. This creates a cycling effect, the

a. CABIN TEMP SEL Rotary Switch.

The CABIN

rate of which can be varied by increasing or decreasing

TEMP SEL rotary switch is labeled COOL and WARM.

the resistance between the temperature selector and the

This switch operates in conjunction with the tempera-

ture controller relay in the heater circuit and with a

thermostat heating winding. Resistance is varied by

cabin thermostat. One set of contacts on the tempera-

turning the CABIN TEMP SEL rotary switch. This

ture controller relay closes to complete a circuit to the

increase or decrease in resistance directly varies the

fuel control solenoid valve. This allows fuel to be

time the heater is allowed to operate before being

delivered to the heater.

automatically cycled.

The second set of contacts on the temperature control-

b. Heater Function Switch.

The heater function

ler relay closes to complete the circuit to the heater

switch is labeled BLWR ONLY, OFF, and HTR ON.

windings in the cabin thermostat. The heater windings

The switch selects the desired feature of the heating and

heat a column of mercury in the thermostat, causing it

ventilating system. When the switch is set to BLWR

to rise. When the mercury column reaches a

34°C

ONLY, the blower forces unheated air into both the

contact, the temperature control relay is shunted, caus-

cockpit and cabin. Further movement of the heater

ing its contacts to open and interrupt the circuit to the

controls is not required. Selecting HTR ON energizes

fuel control solenoid valve. This stops heater operation

the various units of the heater once the HTR START

by shutting off the fuel supply to the heater.

2-50

TM 55-1520-240-10

switch is pressed. The heating and ventilating system is

forward cockpit section heating. When the CABIN AIR

shut down when the switch is set to OFF.

handle is pulled, heated or ventilating air flows through

the ducting to the cabin.

c. HTR START Switch.

When HTR ON is selected

on the heater function switch and the HTR START

2-144. Cabin Heat Controls.

switch is pressed, the heater control circuits are ener-

Fourteen manually adjustable outlets are provided in

gized. The blower starts and purges the heater combus-

the cabin for the comfort of the passengers.

tion chamber of any unburned fuel, while the remainder

of the circuit remains inactive because of a

10 to 15

2-145. Heater Caution.

second time-delay relay. After the time-delay relay is

energized, the ignition assembly is powered and the

NOTE

master fuel solenoid valve opens, allowing fuel to flow to

the heater fuel control unit to complete the start.

Since the HEATER HOT caution will not

extinguish until the temperature in the com-

2-142. Cockpit Air Knob.

bustion chamber is

below 177°C,

it may take

several attempts at restarting the heater

Two cockpit air knobs (15, fig. 2-8 and 9, fig. 2-10) are

before the HEATER HOT caution extin-

on the lower outboard comer of both the pilot and the

guishes.

copilot instrument panels. The knobs are labeled PULL

FOR COCKPIT AIR. Each knob controls a valve on the

heater ducting which regulates the airflow to the cock-

A heater caution capsule labeled HEATER HOT is on

pit.

the master caution panel (fig. 2-51). This caution indi-

cates failure of the automatic temperature control cir-

2-143. Air Control Handles.

cuit. If air temperature in the heater

rises to 177°C

, an

Two air control handles are mounted through a placard

overheat switch deenergizes the automatic temperature

on the right side of the canted console. The placard is

controller relay, shuts off the heating system, except the

labeled AIR CONTROL PULL FOR ON with each

blower, and activates the HEATER HOT caution. The

handle labeled COCKPIT DEFOG OR DEFROST and

heating system will not operate until the blower has

CABIN AIR. By pulling the DEFOG OR DEFROST

lowered the heater temperature to normal and the HTR

handle, heater or ventilating air is directed to the

START switch is pressed. Even though the temperature

cockpit nose enclosure ducting. The airflow is directed

in the combustion chamber has lowered, the HEATER

to the transparent portion of the jettisonable doors and

HOT caution will not extinguish until the HTR START

nose enclosure providing defrosting as well as additional

switch is pressed.

2-51

TM 55-1520-240-10

SECTION Xl ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEMS

2-146. Electrical Power Supply System.

2-147. AC System.

The AC system supplies 115/200-volt three-phase 400-Hz

Alternating current (AC) is the primary source of power

power from No. 1 AC generator to No. 1 three-phase

to operate the electrical and electronic equipment.

AC bus and from No. 2 AC generator to the No.

Three AC generators, two driven by the aft transmission

three-phase AC bus (fig. 2-38). The AC equipment is

and one driven by the APU, produce 115/200-volt

3-phase 400-Hz power. The system develops 28-volt DC

operated by 115-volt single-phase AC and some equip-

through two transformer rectifiers (RECT) one each in

ment by 26-volt AC power supplied through the trans-

the forward section of the left and right fuselage pods.

formers.

DC is also supplied by a 24-volt nickel-cadmium battery.

The AC system is protected from overvoltage, under-

Both 115/200-volt 3- phase AC and 28-volt DC can be

voltage, and underfrequency conditions by generator

supplied by operating the APU or by connecting an AC

control units. The generators will be disconnected from

external power source to the external power receptacles

the AC buses any time the RRPM

drops below 82

to 85

(fig. 2-36). If the APU is running or AC external power

percent for more than 3

to 7 seconds. The AC power

is connected, DC power is supplied by the helicopter

distribution system has four power sources, a contactor

transformer rectifiers (RECT). If only DC external

control circuit, an AC power transfer circuit, and two

power is supplied, AC power is not available on the

AC buses.

helicopter. Circuits are protected by circuit breakers

The No. 1 and No. 2 generator power sources are two

(fig. 2-37). The electrical load is divided between the

main generators driven directly by the aft transmission.

two AC generators (fig. 2-38). Should one generator

The APU generator is driven directly by the APU. The

fail, the other will automatically take over the entire

external power source is an AC power supply connected

load. When the APU is running, its single generator

to the helicopter.

powers the entire load.

No. 1 and No. 2 generators feed their respective buses.

If No. 1 or No. 2 generator fails (or are shut down), the

failed generator is isolated from its bus and the operat-

ing generator feeds both buses. When No. 1 or No. 2 or

both generators are operating, APU generator and

external power are blocked from the AC buses.

When the APU generator is operating and the main

generators are shut down (or rotors turning

below about

84%)

or switched off, the APU generator feeds both

buses. When the APU generator is operating, external

power is blocked from the AC buses. When external

power is applied to the helicopter (GEN APU, GEN

and GEN 2 are OFF), the external power source feeds

both buses.

The generator control unit (GCU) also provides gener-

ator feeder fault protection. If a fault occurs between

the feeder and the airframe, the GCU will disable the

generators. This prevents structural damage to the

airframe when a ground fault occurs.

The permanent magnet generator (PMG) section within

the generator is used to power the main contractors

(relays) in the distribution system. A pickoff coil within

the PMG provides an RPM signal for the rotor tachom-

eter indicators. This tachometer signal is available when-

ever the rotors are turning.

2-148. Generator Control Switches. The generator

control switches are located on the ELEC panel of the

overhead switch panel (fig. 2-39). The three switches are

labeled GEN 1, GEN 2, and GEN APU. The switch

positions are TEST, OFF RESET, and ON.

When the switches are ON, the respective main relay

Figure 2-36. External Power Receptacles

operates, which energizes and connects the generator to

TM 55-1520-240-10

(fig. 2-40). RECT convert 200 VAC power to 28-volt DC

power for use in the DC distribution system.

Cooling air for the RECT is obtained from within the

cabin. The air inlets are located at sta. 176 on the left

and right side of the cabin behind the troop seats. If the

inlets are blocked, the RECT will overheat.

A bus-tie relay is between No. 1 and No. 2 DC buses. If

either RECT fails, the respective RECT failure relay

operates and the bus-tie relay closes automatically to

connect the unpowered bus to the operating RECT. In

addition to No. 1 and No. 2 DC buses, the DC system

includes an essential bus, a switched battery bus, and a

battery bus.

During normal operation, the essential bus and the

Figure 2-39. Electrical Power Panel

switched battery are energized by No. 1 DC bus. If both

DC buses fail or if No. 1 DC bus fails and does not

the buses. At OFF RESET, the generator is deener-

bus-tie to No. 2 DC bus, the essential bus, the switched

gized and disconnected from the bus. This position is

battery bus, and the battery bus will be energized by the

also used to reset a generator. The TEST position is

battert as long as the BATT switch is ON. These buses

provided to allow the generator to be energized but

provide power to emergency, ground maintenance, and

disconnected from the bus to determine whether the

communications components. The battery bus and

AC produced is of proper frequency and voltage.

switched battery bus are energized as long as the battery

2-149. GEN OFF Cautions. Two generator caution

is connected. The hydraulic reservoir level indicators

capsules labeled NO. 1 GEN OFF and NO. 2 GEN

and the emergency APU control circuits and cabin and

OFF are on the master caution panel (fig. 2-51). These

maintenance lights are on these buses.

caution capsules illuminate whenever the generators are

The 24-volt nickel-cadmium battery is located in the left

inoperative. The capsules are controlled by the main

forward electrical compartment. The battery capacity is

generator contractors when the generator control

11 ampere-hours. A battery charger is connected to the

switches are in either ON or OFF RESET. In TEST, the

battery. The battery charger receives power from No. 1

capsules are controlled by the generator control switch

AC bus, rectifies the AC and applies the DC to the

and will extinguish if generator output has the proper

battery to maintain a charge on the battery.

frequency and voltage. Power to operate the generator

capsules is supplied by the DC essential bus through the

Sensors in the battery charger detect battery or battery

LIGHTING CAUTION PNL circuit breaker on the No.

charger overtemperature, short or open circuits or cell

1 PDP.

imbalance. If any of these conditions occur, the battery

charger will stop functioning and activates the BATT

2-150. EXT PWR Caution.

SYS MAL caution capsule on the master caution panel,

CAUTION

External DC power is supplied to the DC buses of the

helicopter by connecting the external DC power source

When external power is used, a visual check

to the DC external power receptacle (fig. 2-36). Appli-

shall be made by the crew to ensure that the

cation of external power operates the DC external

external power unit has been disconnected

power relay which connects the power source to No.

from the helicopter before taxiing.

DC bus. No. 2 DC bus is energized when the bus tie

relay operates. If the polarity of the external power is

reversed, a blocking diode in the circuit prevents the

An external power caution capsule labeled EXT PWR

external power relay from closing.

is on the master caution panel (fig. 2-51). This capsule

illuminates and remains illuminated whenever external

2-152. BATT Switch.

power is connected. The light is controlled by the AC

external power contactor and the DC power relay. The

NOTE

capsule extinguishes when the generators are supplying

The following information applies only if the

current to the buses. Power to operate the external

battery is the only source of power.

power caution capsule is supplied by the DC essential

bus through the LIGHTING CAUTION PNL circuit

breaker on the No. 1 PDP.

The BATT (battery) switch is located on the ELEC

panel of the overhead switch panel (fig. 2-39) The

2-151. DC System.

two-position switch is labeled ON and OFF. When the

The direct current (DC) power supply system supplies

switch is at ON, the essential, switched battery, and

28-volt DC from the No. 1 transformer rectifier (RECT)

battery buses are energized. Regardless of the battery

to No. 1 DC bus and the No. 2 RECT to No. 2 DC bus

switch position, the switched battery and battery buses

2-57/(2-58 blank)

TM 55-1520-240-10

are powered directly by the battery. To prevent exten-

LIGHTING CAUTION PNL circuit breaker on the No.

sive discharging of the battery while making extended

1 PDP.

ground checks of equipment, use an external electrical

power source or operate the APU generator.

2-154. BATT SYS MAL Caution.

2-153. RECT OFF Cautions. Two RECT caution cap-

A battery system malfunction caution capsule labeled

sules labeled NO. 1 RECT OFF and NO. 2 RECT OFF

BATT SYS MAL is on the master caution panel (fig.

are on the master caution panel (fig. 2-51). These

2-51). This caution illuminates when the battery charger

has stopped charging the battery. This can be caused by

caution are controlled by the reverse-current cutouts.

an overheated battery or battery charger, battery cell

Whenever one of the RECT fails, either through a fault

imbalance, or an output short or open circuit. Power to

in the RECT or a bus fault, the respective caution

operate the capsule is supplied by the DC essential bus

illuminates. Power to operate the transformer rectifier

through the LIGHTING CAUTION PNL circuit

capsules is supplied by the DC essential bus through the

breaker on the No. 1 PDP.

2-61

TM 55-1520-240-10

SECTION XII AUXILIARY POWER UNIT

2-155. General.

or white. A label on the ESU explains the various BITE

indications and their meaning.

The gas turbine auxiliary power unit T62-T-2B (APU)

(fig. 2-41) is mounted in the aft cabin above the ramp.

2-157. APU Switch. The APU switch is on the ELEC

The basic components of the APU are the gas turbine

panel of the overhead switch panel (fig. 2-39). It is a

engine, hydraulic motor-pump, fuel control, accessory

three-position switch labeled OFF, RUN, and START.

drive, and AC generator. An APU ELECTRONIC

The switch is spring loaded from START to RUN.

SEQUENCING UNIT (ESU) which monitors APU

Normally, power to operate the APU is supplied by the

operation is on the left side of the cabin above the ramp.

DC essential bus through the APU CONT NORM

The ESU is also labeled APU CONTROL BOX.

circuit breaker on the No. 1 PDP. Emergency power to

The motor-pump on the APU pressurizes the utility and

operate the APU is from the battery bus through the

hydraulic system for main engine starting and ground

APU CONT EMERG circuit breaker on the No. 1 PDP.

checks. The APU also drives an AC generator which

supplies power to No. 1 and No. 2 electrical systems.

2-158. APU ON Caution. The APU ON caution cap-

Refer to Section VI for further information on the

sule is on the master caution panel (fig. 2-51). Normally,

hydraulic systems. The APU oil supply is integral and

the APU is intended for ground operation only. It is not

contained within the sump of the accessory drive assem-

intended for operation during flight. If the caution

bly. The APU receives fuel from the left main fuel tank

remains illuminated following take-off, it alerts the pilot

through a booster pump, a manual fuel shutoff valve,

to shut down the APU. When the caution is illuminated,

and a solenoid valve.

it indicates the APU is up to speed and the exhaust gas

temperature is normal, It does not necessarily indicate

2-156. ELECTRONIC SEQUENCING UNIT.

that APU hydraulic pump or generator output is nor-

The ESU is mounted on the left side of the cabin above

mal. If the rotors are not turning, check the UTIL HYD

the ramp. The unit monitors APU starting and opera-

SYS and RECT OFF cautions to evaluate output of the

tion. In addition, it monitors APU speed and exhaust

APU hydraulic pump and generator. The APU ON

gas temperature. The unit continuously compares these

caution is controlled by the ESU.

parameters with limits programmed into ESU circuits.

If a limit is exceeded, the ESU will automatically shut

2-159. EMERGENCY APU FLUID SHUT OFF

down the APU.

VALVE.

The EMERGENCY APU FLUID SHUT OFF VALVE

NOTE

is in the fuel supply line to the APU (fig. 2-42). It is

The BITE indicators indicate engine condi-

located inside the aft cabin above and to the left of the

tion only. They will not indicate a defective

ramp interphone station. The valve can also be reached

hydraulic motor-pump or generator.

from the outside through an access door labeled AC-

CESS APU EMER FLUID SHUT OFF. The knob on

Four magnetic built-in-test-equipment (BITE) indica-

the valve has an OPEN and CLOSE position. Placing

tors are on the ESU. These indicators are either black

the knob to CLOSE shuts off fuel to the APU.

2-62

55-1520-240-10

SECTION XIII

LIGHTING (NVG)

2-160. Position Lights.

NOTE

The crew chief must inform the pilot w

Three position lights (1, 4, and 6, fig. 2-43) are installed on

AFT POS LIGHT switch has been changed

the helicopter. On the right side of the fuselage is a green

the OFF position.

light (1); on the left, red (6); and on the vertical panel of the

aft pylon, white (4). Power to operate the position lights is

supplied by the No. 2 DC bus through thb. AFT POS LIGHTT

he AFT POS LIGHT

switch is located in the cabin at sta. 5

circuit breaker on the No. 2 PDP.

TENANCE PANEL (fig. 2-45). The guarded t

2-161. Position Light Switches.

switch is labeled OFF and ON. It allows

light to be turned off during aided (NVG) o

POSN Light

The POSN (position) sduringiunaided night operations.

located on the EXT LTG (exterior lighting) panel on the left

side of the overhead switch panel (fi2-162.Formation)Lights.he three-

position switch is labeled DIM, OFF, anThere are five elec

troluminescent panels

intensity of the position lights. Whenformation operations (

2 and 5, fig. 2-43

position light system is deenergized. compatible formation lights for NVG forma

(9, fig. 2-43).

A50277

1. Right position light (Gre6.

Left position light (Red)

2. Formation lights (3)

7. Bottom anticollision light

3. Top anticollision light

8. Landing-search lights (2)

4. Position light (White)

9. NVG formation lights (8)

5. Formation lights (2)

Figure 2-43. Exterior Lights

Change

2-65

55-1520-240-10

a.ElectroluminesceThree panels whichnormal operations use the pilot searchlight

an equilateral triangle are aft of tlight. For NVG operations use the copilot s

panels are on the top of the aft pylonlandingtlight.icollision

light. Power to operate and control the electroluminescent

formation lights is supplied by th2-167.ISRCHLTCONTRSwitch.M

Two SRCHLT

circuit breaker on the No. 1 PDP.

CONTR (searchlight control) switches are on t

switch panel (fig. 2-44). The PLT SRCHLT CONT

b.NVG Formation

There is an NVG formation

is on the PLT LTG panel. The CPLT SRCHLT C

light on each side of the forward pylon, two NVG formation

switch is on the CPLT LTG panel. Each two-pos

lights on each side of the fuselage, two NVG formation

is labeled RET and ON.

lights on the aft pylon, one aft of the anticollision light, and

one on the vertical panel at the rear oWhen the SRCHLT CONTR

switch is placed to O

he THRUST CONTR lever be

to operate and control the NVG formatioSLT-FIL switch on t

by the No. 1 DC bus through the LIGHToperational.

If the searchlight is at any angl

circuit breaker on the No. I PDP.

the SRCHLT CONTR switch is placed to RET, the

light will automatically rotate to point forw

2-163. FORM Light Switches.

The FORM (formaretract flush with the fuselage. Power is supp

tion) light select and control switch1 and No. 2 DC bus through the LIGHTING SL

EXT LTG panel on the left side of tcircuit breakers on

the No. 1 and No. 2 PDP.

panel (fig. 2-12).

2-168. SLT-FIL Switch.

a.FORM Light SelecA twoŧposition toggle

switch labeled NVG and NORM. In the NORM position,

CAUTION

the five electroluminescent panels may be controlled by the

FORM light rotary control switch for norThe copilot landing searchlight emits

tion operations. In the NVG position,ible infrared rays which may be hazardou

‘formation lights may be controlled bypersonnel looking directly at the light

rotary control switch for NVG night formrange or touching it. Ensure that the

landing SLT-FIL switch Is OFF and the

b.FORM Light Contro

A rotary control switch

ted when it is not in use.

labeled OFF, DIM, and BRT with three efully retrac

incremental markings between DIM and BRT. It adjusts the

intensity of the formation lights selecA SLT-FIL (searchlig

ht filament) switch is

select switch. When the rotary contropilot and copilot THRUST CONT lever switch b

formation light system is deenergized.

2-26.) Each switch is labeled ON and OFF. T

turn on the landing-searchlight lamp, before

2-164. Anticollision Lights.

sion. Power to operate the landing search

Two red strobe anticollision lights asupplied by the No. 1

and No. 2 DC bus t

and 7, fig. 2-43). One is on top of theLIGHTING SLT FIL circuit breakers on the No

is on the fuselage underside. Power to2 PDP.e the anticol-

lision lights is supplied by the No. 2 DC bus through the

2-169. SEARCH LIGHT Position Switch.

LIGHTING ANTI COL TOP and BOT circuit breakers on

the No. 2 PDP.

CAUTION

2-165. ANTI COL Light Switches.

Two ANTI COL

Do not confuse the SEARCH LIGHT posi-

TOP and BOT toggle switches are on the EXT LTG panel

tion switch with the two engine beep

on the left side of the overhead switch panel (fig. 2-12).

switches.

Each two-position switch is labeled OFF and ON. When the

anticollision light switch is ON, the lights are energized.

When the switch is placed to OFF, theA five-position momentary SEARCH LIGHT swit

are deenergized.

each THRUST CONTR lever switch bracket (fig

is labeled L (left), EXTEND, R (right), and

2-166. Landing Searchlights.

switch is spring-loaded to center off posit

Two controllable landing-searchlights aWhen the SRCHLT CO

NTR and SLT-FIL switches ar

bottom of the fuselage (8, fig. 2-43).the searchlight can b

e controlled up and do

the pilot THRUST CONT lever and theright with thehe

SEARCH LIGHT position switch

copilot THRUST CONT lever. The copilotoperate the searc

hlight position switch is

with an infrared (IR) filter for NVGNo. 1 and

No. 2 DC bus through the LIGHT

CONT circuit brea

1and No. 2 PDP.

Each light is operated independently by a SRCHLT CONTR

switch, SLT-FIL (searchlight filament), and position switch.

170. Overhead Switch Panel Lights.

They may be extended and stopped at any angle up to 90°

in a vertical plane and rotated 360° abThe overhead switch panel has integral light

long as the searchlight position switoperate and control the overhead panel lights

2-66

Change

TM 55-1520-240-10

The rotary control switch is labeled OFF, DIM, and

2-176. Canted and Center Console Lights.

BRT. It adjusts the light level from DIM to BRT. When

NOTE

the rotart control switch is OFF, the overhead switch

The console lights are incompatible with

panel light system is deenergized.

NVG. During NVG operations, turn the

console lights off and light the console with

2-172. Pilot and Copilot Instrument Panel Lights.

utility lights or floodlights.

All flight instruments and placards on both pilot and

copilot instrument panels receive lighting. The HSI,

Lighting is provided for all control panels on the canted

attitude indicator (VGI), radar altimeter, and turn and

and center console. Power to operate and control the

slip indicator for both pilot and copilot have integral

console lights is supplied by the No. 1 AC bus through

lighting. The remaining instruments are externally lit by

the LIGHTING CONSOLE circuit breaker on the No.

lighting posts adjacent to the instruments. Power to

1 PDP.

operate and control the pilot flight instrument lights is

supplied by the No. 2 AC bus through the LIGHTING

2-177. LTG Panel. The LTG panel is located at the

PILOT INSTR circuit breaker on the No. 2 PDP. Power

rear of the overhead switch panel (fig. 2-44). It consists

to operate and control the copilot flight instrument

of the CTR CSL and STICK POSN IND control

switches.

lights is supplied by the No. 1 AC bus through the

LIGHTING COPLT INST circuit breaker on the No.

a. CTR CSL Switch.

Rotary control switch labeled

PDP.

OFF, DIM, and BRT. It adjusts the light level on the

canted and center consoles from DIM to BRT. When

2-173. PLT and CPLT INST Switches. The PLT

the rotary control switch is OFF, the canted and center

INST (pilot instrument) control switch is located on the

consoles light system is deenergized.

PLT LTG panel on the right side of the overhead switch

b. Stick POSN IND Switch.

Rotary control switch

panel (fig. 2-44). The CPLT INST (copilot instrument)

labeled OFF, DIM, and BRT. It adjusts the light level

control switch is located on the CPLT LTG panel on the

on the stick position indicator from DIM to BRT. When

left side of the overhead switch panel. The rotary

the rotary control switch is OFF, the stick position

control switches are labeled OFF, DIM, and BRT. They

indicator light system is deenergized.

adjust the light level from DIM to BRT. When the

rotary control switch is OFF, the respective instrument

2-178. Dome Lights.

panel light system is deenergized.

When the PLT INST rotary control switch is placed out

of the OFF detent, the following lighting is dimmed:

If the white dome light is turned on during

a. Troop warning jump lights on the overhead switch

NVG operations, the effectiveness of the NVG

panel and on the emergency troop alarm and jump

may be severely impaired and a hazardous

lights boxes on the cargo compartment.

situation may be created due to sudden loss

of pilot visual references. Do not turn on the

b. The legend on pushbutton switches on the heading

white dome lights during NVG operations.

and altitude section of the AFCS panel.

Two cockpit dome lights are attached to the overhead

c. The legend on pushbutton switches on the pilot

structure adjacent to the overhead switch panel (fig.

and copilot HSI MODE SELECT panels.

2-44). Each dome contains a white lamp and a blue

d. The legend on STATUS pushbutton switch on the

NVG filtered lamp which can be selected individually.

countermeasure set AN/ALQ-156 control panel.

Power to operate and control the dome lights is sup-

plied by the DC essential bus through the LIGHTING

2-174. Center Instrument Panel Lights.

COCKPIT DOME circuit breaker on the No. 2 PDP.

The center instrument panel as well as the fire warning

2-179. DOME Switch. The DOME switch is located

panel are lighted. Power to operate and control the

on the INTR LTG panel at the right rear of the

center instrument panel lights is supplied by the No.

overhead switch panel (fig. 2-44). The three-position

AC bus through the LIGHTING CTR INSTR circuit

positive-locking switch is labeled WHT, OFF, and NVG.

breaker on the No. 2 PDP.

It selects the function of the dome light. The center

position lever locking switch prevents inadvertent white

2-175. CTR INST Switch. The CTR INST control

light activation during NVG operations.

switch is located on the PLT LTG panel on the right

When the DOME switch is placed to WI-IT, the master

side of the overhead switch panel (fig. 2-44). The rotary

caution panel cannot be dimmed. If WHT is selected

control switch is labeled OFF, DIM, and BRT. It adjusts

while the caution panel is operating on DIM, the

the light level from DIM to BRT. When the rotary

caution lights will automatically switch to BRT mode.

control switch is OFF, the respective instrument panel

During NVG operations, the DOME switch should only

light system is deenergized.

be placed to NVG.

2-68

Change 1

TM 55-1520-240-10

2-180. Pilot and Copilot Utility Lights.

an NVG blue lamp and a red lamp which can be

selected individually. Power to operate and control the

Two utility lights, connected to individual flexible cords,

cabin and ramp lights is supplied by the switched battery

are mounted in two retaining sockets on either side of

bus through the LIGHTING CABIN & RAMP circuit

the overhead switch panel above the pilot and copilot

breaker on the No. 1 PDP.

(fig. 2-44). The lights are detachable and can be moved

about to take care of special lighting situations. Each

2-185. CABIN AND RAMP LIGHTS Switches. The

utility light has a rheostat switch as an integral part of its

CABIN AND RAMP LIGHTS switches are located on

assembly. This switch, located on the aft part of the

a control panel below the ramp control lever. The

light, regulates the intensity of the light from OFF to

control panel consists of a select switch and a CON-

BRT. A white button on the light housing, opposite the

TROL rotary switch.

switch, is used for flashing the light. By selecting the

color desired on the barrel of the light, blue or white

a. Select Switch. Three-position toggle switch labeled

light will be emitted. Power to operate the utility light is

RED, OFF, and NVG. It is used to select the appro-

supplied by the No. 2 DC bus through the LIGHTING

priate cabin and ramp lights. When placed to OFF,

COCKPIT DOME circuit breaker on the No. 2 PDP.

cabin and ramp lights are deenergized.

b. CONTROL Switch.

Rotary switch labeled DIM

2-181. Floodlights.

and BRT. It adjusts the cabin and ramp RED or NVG

Eight floodlights provide a secondary source of light

light level from DIM to BRT.

(fig. 2-44). Six are under the glareshield and two on the

cockpit bulkhead. The six floodlights under the

2-186. Emergency Exit Lighting.

glareshield light the pilot, center, and copilot instru-

ment panel. The two overhead floodlights light the

Three emergency exit lights are in the cargo compart-

overhead switch panel. Power to. operate and control

ment close to each of the three primary emergency exits

the floodlights is supplied by the DC essential bus

(fig. 2-46). They are located by the main cabin door, the

through the LIGHTING INSTR FLOOD circuit

emergency exit opposite the main cabin door, and the

breaker on the No. 2 PDP.

ramp emergency exit. The lights come on whenever a

loss of power on the switched battery bus occurs or

2-182. FLOOD Switches. The FLOOD switches are

during a landing when 3 to 4g’s are exceeded as sensed

located on the INTR LTG panel at the right rear of the

by an inertia switch.

overhead switch panel (fig. 2-44). They consist of two

floodlight selection switches and a rotary control switch.

The emergency exit lights system is controlled by the

The floodlight selection switches are labeled INST and

EMER EXIT switch on the INTR LTG panel of the

OVHD. Each switch has an OFF and ON position. The

overhead switch panel. The lights may also be used as

rotary control switch is labeled OFF, DIM, and BRT.

portable lamps by removing them from their housing

and by rotating the handle, marked PULL EMER-

a. INST and OVHD Floodlights Selection Switches.

GENCY LIGHT, 45° from its normal position. Power to

Each switch is labeled for the area the floodlights will

operate the emergency exit lights is supplied by two,

light. By placing either switch ON, the associated flood-

internal, 1.25 volt, nickel-cadmium batteries. Power to

lights will light when the floodlight rotary control switch

operate and control the charging, monitoring, and test

is turned toward BRT. Placing the switch to OFF

circuit is supplied by the switched battery bus through

deenergizes the floodlight circuit.

the LIGHTING EMER EXIT circuit breaker on the

b. Floodlight Rotary Control Switch.

The rotary con-

No. 1 PDP.

trol switch is used to adjust the floodlights from DIM to

BRT once the respective floodlight selection switch is

2-187. EMER EXIT Switch.

placed to ON. When the rotary control switch is OFF,

the floodlights will be deenergized.

CAUTION

2-183. Emergency Floodlights.

If the EMER EXIT switch is left in ARM or

If the pilot flight instrument lights have been turned on,

DISARM with the helicopter shutdown and

loss of electrical power will cause the floodlights to

the battery connected, the charging circuit of

automatically come on. All floodlights will function

the emergency exit light system will dis-

automatically in the BRT mode. Simultaneous dimming

charge the helicopter battery.

control of all floodlights can be regained by setting the

FLOOD INST and OVHD selection switches to ON,

turning the floodlight rotary control switch to BRT on

The EMER EXIT switch is located on the INTR LTG

the INTR LTG panel, and turning the PLT INST rotary

panel of the overhead switch panel (fig. 2-44). The

control switch to OFF. Floodlight intensity can be

three-position switch is labeled DISARM, TEST, and

ARM. When the switch is placed to ARM, the emer-

controlled by the floodlight rotary control switch on the

INTR LTG panel.

gency exit lights stay off, the batteries are charging, and

the charge indicator lights come on. The circuit moni-

2-184. Cabin and Ramp Lights.

tors electrical failures and landings in

excess of 3 to 4g’s.

Five cabin and ramp lights are in the cabin, attached to

The light from the charge indicator lamps can be seen

the overhead structure (fig. 2-45). Each light contains

emitting through two pin holes at the base of the main

2-69

TM 55-1520-240-10

SECTION XIV FLIGHT INSTRUMENTS

2-190. General.

stress when in the red-and-yellow striped band. This can

be accomplished by lowering THRUST CONT lever,

The following paragraphs contain information on the

reducing airspeed, releasing back pressure on the cyclic

flight instruments. Information on the navigation instru-

stick, or by reducing the severity of the maneuver.

ments will be found in Chapter 3, Avionics. All other

instruments directly related to one of the helicopter

2-193. CGI TEST Switch. The CGI TEST switch is

systems are found under the appropriate system head-

on the pilot instrument panel, on the left side of the

ing in this chapter. Refer to fig. 2-6, 2-7, 2-8, 2-9, 2-10,

indicator (fig. 2-47). It is a three-position switch spring

and 2-12 for illustrations of the instrument panels,

loaded to center off position labeled FWD and AFT.

canted and center consoles, and overhead switch panel.

NOTE

2-191. Cruise Guide Indicator System.

Do not test the cruise guide system with

The cruise guide indicator (CGI) system gives the pilot

rotors turning. False indications will result.

a visual indication of actual loads imposed on critical

components of the helicopter dynamic system, The

When the switch is placed from center position to each

system allows the pilot to achieve maximum helicopter

utilization under various conditions of payload, altitude,

test position, the pointer on the indicator should indi-

airspeed, ambient temperature, and center-of-gravity.

cate within the white test band. The white test band

indicates proper system operation.

The system consists of strain gages bonded to fixed links

in the forward and aft rotor controls, an indicator, a

When the test function is activated, circuits from the

signal processor unit in the aft pylon, a signal condi-

strain gages to the indicator are tested. However, sep-

tioner unit in the forward pylon, and interconnecting

aration of the bond of the strain gage to a link will not

wiring. The system measures alternating stress loads at

be detected by the test function. The narrow white line

each rotor and displays the larger of the two signals.

towards the high end of the striped red-and-yellow band

Power to operate the cruise guide indicator system is

is used for calibrating the indicator during bench test.

supplied by No. 2 DC bus through the CRUISE GUIDE

circuit breaker on the No. 2 PDP.

2-194. Airspeed Indicator.

There are two airspeed indicators located on the upper

2-192. CRUISE GUIDE Indicator.

The CRUISE

left portion of the copilot and pilot instrument panel

GUIDE indicator is on the pilot instrument panel (fig.

(fig. 2-8 and 2-10). The difference between dynamic

2-47). Three bands are displayed on the dial face of the

pressure and static pressure as measured by the pitot

indicator. These bands are colored green, yellow, and

static system is introduced into these instruments. Indi-

striped red-and-yellow, Refer to fig. 5-1 for limitations.

cated airspeed is shown in knots.

Immediate corrective action must be taken to reduce

2-195. Altimeter.

An AIMS altimeter is provided for the pilot (fig. 2-48).

In the term AIMS, the A stands for Air Traffic Control

Radar Beacon System (ATCRBS), the I stands for

identification friend or foe (IFF), the M represents the

Mark XII identification system, and the S means system.

A pneumatic counter-drum-pointer type altimeter is

installed for the copilot. The pilot’s altimeter is a

pneumatic counter-drum-pointer type which is a self-

contained unit consisting of a precision pressure altim-

eter combined with an altitude encoder.

Simultaneously, the display indicates and the encoder

transmits through the transponder the altitude of the

helicopter. Altitude is displayed on the altimeter by a

10,000-foot counter, and a 100-foot drum. A single

pointer indicates hundreds of feet on a circular scale

with 50-foot center markings. Below 10,000 feet, a

diagonal warning symbol will appear on the 10,000 foot

counter. Power to operate the AIMS altimeter is sup-

plied by the No. 2 DC bus through the NAV AIMS ALT

circuit breaker on the No. 2 PDP.

A barometric pressure setting knob is provided to insert

Figure 2-47. Cruise Guide Indicator

the desired altimeter setting in inches of Hg. A vibrator

2-72

TM 55-1520-240-10

b. Check that the pilot’s altimeter indicates within

±70 feet of field elevation. If the altimeter error is

greater than

±70 feet, do not use the altimeter for IFR

flight.

2-197. In Flight Operation - Altimeter.

Operate the AIMS altimeter encoding function as fol-

lows:

a. Be sure the IFF set is on and set to the proper

code.

b. Be sure the altimeter is set to the local altimeter

setting.

c. Set the M-C (mode c) switch on the IFF control

panel to ON.

d. Check that the red CODE OFF flag is not visible in

the pilot’s altimeter.

The copilot’s altimeter is a pneumatic counter-drum-

pointer type which displays altitude in the same manner

as the pilot’s altimeter. It also incorporates a barometric

pressure setting knob and an internal vibrator powered

Figure 2-48. AIMS Altimeter

by the No. 2 DC bus. A

minimum of 1

minute of vibrator

operation is required before setting or checking the

altimeter.

and requires a minimum of 1 minute warmup before

checking or setting the altimeter. If DC power to the

At ambient pressure,

both altimeters should agree

altitude encoder is lost, a warning flag placarded CODE

within

±70 feet of the field elevation when the proper

OFF appears in the upper left instrument dial.

barometric pressure setting is set in the altimeter. If the

internal vibrator of either altimeter becomes inopera-

The flag indicates that the altitude encoder is inopera-

tive due to DC power failure, the pointer drum may

tive and that the system is not reporting altitude to

momentarily hang up when passing from

9 through

ground stations. The CODE OFF flag monitors only the

encoder function of the altimeter. It does not indicate

(climbing) or from

0 to 9 (descending). This will cause a

transponder condition. The AIMS altitude reporting

lag of magnitude which will depend on the vertical

function can be inoperative without the CODE OFF

velocity of the aircraft and the friction in the altimeter.

flag showing, as in case of transponder failure or

2-198. Radar Altimeter (AN/APN-209A). Radar al-

improper control settings. It is also possible to get a

timeters are provided for the pilot and copilot (fig.

good Mode C test on the transponder control with the

2-49). The altimeters provide a continuous indication of

CODE OFF flag showing. Display of the CODE OFF

the height of the helicopter above the surface from

0 to

flag only indicates an encoder power failure or a CODE

1,500 feet. Altimeter indications are reliable with pitch

OFF flag failure. In this event, check that DC power is

and roll attitude up to 45°.

available and that the circuit breakers are in. If the flag

is still visible, radio contact should be made with a

Altitude is displayed by a dial, pointer, and by a digital

ground radar site to determine whether the AIMS

display. Each altimeter has HI and LO caution lights.

altitude reporting function is operative. The remainder

The caution lights on each altimeter can be set inde-

of the flight should be conducted accordingly.

pendently of the other altimeter. The caution lights are

set by rotating the LO SET and HI SET knobs until the

2-196. Preflight Operation - Altimeter.

L index and H index on the perimeter of the altimeter

If the AIMS altimeter encoding function is to be used

are at the desired altitudes.

during a flight, perform the following steps before

takeoff

When the helicopter descends below the low index

setting or rises above the high index setting, the corre-

CAUTION

sponding HI or LO caution light will illuminate. If

helicopter altitude exceeds

1,500

feet, pitch or roll angle

If the baroset knob binds or sticks, do not

exceeds 45°,

or the system is unreliable, the following

use excessive force to set the altimeter. Ex-

will occur. The OFF flag will appear, the pointer will

cessive force can damage altimeter gears,

move through

1,500 feet behind the dial mask, and the

resulting in altimeter error. Settings can

digital display and LO and HI caution lights will extin-

sometimes be made by backing off the knob

guish.

and turning at a slower rate.

If power to the system is lost, the following will occur.

a. Set the pilot’s altimeter to the field barometric

The OFF flag will appear, the digital display and

setting.

caution lights will go out, and the pointer will remain at

2-73

TM 55-1520-240-10

Figure 2-49. Radar Altimeter (AN/APN-209)

the last valid indication when power was lost. The

CONTROLS/

altimeters have a self-test feature. Pressing the PUSH-

INDICATOR

FUNCTION

TO-TEST knob will cause the pointer and digital dis-

LO set index

Indicates altitude trip

play to indicate between 900

and 1,100 feet. If LO set

point for LO caution light.

and HI set are indexed

below 900

feet, the LO caution

HI set index

Indicates altitude trip

light goes out, and the HI caution light comes on. Power

point for HI caution light.

to operate the radar altimeter is supplied by the No. 2

Indicator pointer

Indicates absolute altitude

DC bus through the NAV RAD ALT circuit breaker on

from 0 to 1,500 feet.

the No. 2 PDP.

Digital indicator

Provides direct reading

four digit indication of

2-199. Controls and Function, Radar Altimeter

absolute altitude from

(AN/APN-209A).

to 1,500 feet.

LO caution light

Comes on when helicop-

CONTROLS/

INDICATOR

FUNCTION

ter descends below alti-

tude on LO set index.

LO SET knob

Either pilot’s LO SET

HI caution light

Light comes on when heli-

knob applies power to

copter rises above altitude

altimeter system. LO set

on HI set index.

index on both altimeters

OFF flag

Flag is displayed when

can be set independently.

power is removed from

Both LO set indices must

set, when indications are

be masked to turn the set

unreliable, or when alti-

off.

tude exceeds approxi-

HI SET knob

Sets position of HI set

mately 1,500 feet.

index and tests altimeter

system when pressed.

2-74

TECHNICAL MANUAL MAINTENANCE MANUAL FOR ARMY CH-47D HELICOPTER (1992) - page 3