UH-3H and UH-3H EXECUTIVE TRANSPORT. FLIGHT MANUAL (2000)

NAVAIR 01-230HLH-1

Figure 2-6. Overhead Engine Controls

extreme caution as it has a positive influence on fuel flow, and

2.1.6 Engine Instruments

misuse can cause engine overspeed or overtemperature. The

lever is mechanically connected to a cam within the fuel control

2.1.6.1 Power Turbine Inlet Temperature Indicators.

that contacts the fuel metering valve. The initial position of the

Two of these indicators (Figure 2 - 7) marked EXH TEMP

fuel metering valve is dependent upon the automatic features of

indicate engine power turbine inlet air temperatures in degrees

the control as established by the setting of the speed selector.

celsius and are on the instrument panel. The indicators operate

The cam, when actuated by advancing the manual throttle,

from thermocouples forward of the power turbine in the

contacts the fuel metering valve. Once contact is established,

second-stage turbine casing on each engine. When the

further advancement of the manual throttle will manually

dissimilar metals of the thermocouples in the engine are heated,

control fuel flow that in turn regulates engine power output.

an electromotive force (independent of the helicopter electrical

The manual throttle is unable to reduce the position of the

system) is created and a resulting current flow through a known

metering valve below that called for by the speed selectors.

resistance of the thermocouple circuit deflects the indicator

Control below this point will depend upon the type of

needle that is read in degrees celsius. The pilot has no direct

malfunction encountered. In all instances of manual throttle

control for regulating the power turbine inlet temperatures;

operation, it must be remembered that the speed selectors

however, limited control for lowering these temperatures can be

should not be retarded beyond GRD IDLE.

indirectly achieved by reducing collective pitch or power

demand.

2.1.6.2 Oil Pressure Indicators. Two of these indicators

WARNING

(Figure 2-7), one for each engine, are on the instrument panel

and indicate oil pressure in pounds per square inch. The

indicators are powered by 26 vac and each is protected by

With manual throttle actuated, resistance may

circuit breakers.

occur in the speed selector during conditions

requiring movement at or below the minimum

2.1.6.3 Oil Temperature Indicators. Two of these

governing range marking on the throttle

indicators (Figure 2 - 7), one for each engine, are on the

quadrant. Attempts to retard speed selector

instrument panel and indicate engine oil temperature in

beyond the point at which this resistance occurs

degrees celsius. The engine oil temperature bulb in the

may result in inadvertent engine shutdown.

bottom of the tank transmits indications to the respective

temperature indicators. The indicators are powered by 28

The fuel stopcock is downstream of the metering valve

vdc and protected by circuit breakers marked 1 ENG 2

and is actuated by the speed selectors. Placing the speed

under the general heading OIL TEMP on the center circuit

selector to SHUTOFF will stop engine fuel flow regardless of

breaker panel.

manual throttle control lever position.

2-15

ORIGINAL

NAVAIR 01-230HLH-1

2.1.7 Starter System. The system consists of a starter,

2.1.7.2 Start Bleed Valve. The start bleed valves, one on

starter relay, start bleed valve, fuel shutoff valve, starter buttons,

each engine, operate automatically during the start cycle and

mode selector switch (manual or normal modes only), an abort

require no specific pilot action. The function of the valve is to

switch, start bleed valve actuating switches, and an emergency

raise the compressor stall line during the start cycle in order to

start switch. The system operates on dc power and is protected

increase the reliability of the normal start system. This is done

by circuit breakers marked STARTER 1 ENG 2 on the center

by bleeding compressor discharge air during start. The valve

circuit breaker panel. The engine starting system (FO-3)

closes at the point where the starter, ignition circuit, and valve

consists of three modes of operation: normal, manual, or

circuit are deenergized simultaneously. The valve remains fully

emergency. The normal mode provides a completely automated

closed during all regimes of engine operation except during

start that includes automatic starter dropout and increased start

starting.

reliability through a start bleed valve. The manual mode

provides an alternate means of starting, using external electrical

2.1.7.3 Starter Buttons. Two starter buttons, one for each

power or the battery. In this mode, the starter relay dropout

engine, are above the speed selectors. The starter is energized

feature is bypassed. This permits the starter to motor the engine

by holding the speed selectors at SHUTOFF and momentarily

continuously until the starter abort switch is actuated by pulling

depressing the starter button. This energizes the starter relay and

down on the speed lever. An electrically operated solenoid

completes the circuit to the starter. When using the normal

valve aids in controlling overtemperature conditions. When the

starting mode, after the engine fires and the electrical power

valve is actuated by the pushbutton-type switch mounted on

load to the starter decreases, the starter relay will automatically

the cyclic stick, auxiliary start fuel flow is blocked from

drop out, deenergizing the starter. In the manual mode, when

entering the fuel flow divider. A series of safety interlocks in

Ng reaches about 45 percent, the respective speed selector must

the normal or manual control circuit to each engine starter

be pulled down to actuate the abort switch that in turn drops out

prevents the starter relay from closing should an unsafe

the starter.

condition exist.

2.1.7.4 Starter Abort Switch. A starter abort switch is in

For engine starts during flight or for starting the No. 1

each speed selector. The switch is actuated by pulling down on

engine when the tail pylon is folded, an emergency start

the speed selector. This action breaks electrical circuit

switch provides a means of bypassing all safety interlocks,

continuity to the ignition system and the starter relay.

thus allowing either a manual or normal start. Before

starting, check that the

10-ampere blade fold circuit

2.1.7.5. Mode Selector Switch. This switch with marked

breaker marked BLADE FOLD on the center circuit

positions MANUAL and NORMAL under the general heading

breaker panel is in.

START MODE is on the overhead dome light panel. When

the switch is placed to NORMAL, the automatic dropout

function of the starter relay is energized, allowing the starter to

motor the engine to about 45-percent N_g. When the switch is

placed to MANUAL, the automatic dropout feature of the

starter relay is bypassed, allowing the starter to remain engaged

until the abort switch is actuated. The switch operates on 28-

vdc power.

2.1.7.6 Auxiliary Start Fuel Shutoff Valves. Two of these

valves, one for each engine, are installed in the engine

compartment between the engine fuel control and flow divider.

When the valve is actuated during the start of either engine, the

flow of auxiliary bypass starting fuel is blocked. This blockage

decreases the total amount of fuel flow during starting, thus

diminishing the possibility of an overtemperature condition

2.1.7.1 Starter Dropout. Starter operation and dropout

because of excessive fuel flow. The valves operate on dc power

may be monitored by noting the magnetic compass heading

and are protected by the main starting circuit breakers.

before engine start. When the starter is energized, the compass

will swing to a new heading. When the normal start circuit is

2.1.7.7

Auxiliary Start Fuel Shutoff Valve Switch.

used, the starter will drop out at 45-percent N_g. The compass

This pushbutton-type switch marked ENG ST is on each cyclic

should then swing back to its original heading, signifying the

stick. In addition to pressing the switch, the starter relay for the

starter has dropped out. When the manual start circuit is used,

engine to be started must be closed before the valve will

the starter will drop out only when the abort switch is actuated.

operate. Either the pilot or copilot switch will control the

operation of both valves. The switch operates on dc power.

2-18

ORIGINAL

NAVAIR 01-230HLH-1

2.1.7.8 Emergency Start Switch. Two switches (Figure 2-

at NORM. When the switch is at NORM with the starter

8) marked EMER START 1 ENG 2 are on the overhead

engaged, the ignition unit is energized. Holding the switch at

control panel to the right of the engine speed selectors. The

the spring-loaded TEST position energizes the ignition unit

switches have two marked positions, ON and OFF. Normally

only. TEST is used (for ground operation only) without the

the switches remain OFF and starting is done through the

starter to test the ignition circuit. A clicking noise can be heard

normal control circuit. When the switches are placed ON, the

when the switch is placed to TEST. When the switch is OFF,

normal control circuit with its safety interlocks is bypassed.

the ignition unit is deenergized. OFF is used to motor the

Power from the primary dc bus is fed directly to the starter

engine, using the starter without ignition. The ignition switches

button and a normal starting procedure is followed. Placing

are powered by dc current.

either engine emergency starter switch ON permits starting of

that engine in flight should any of the safety interlocks fail to

2.1.9 Torque Sensing System.

This system

remain in the closed position when deenergized. The

determines input torque at the main gearbox and transmits

emergency starter switches also permit intentional interlock

this information to torquemeter indicators. Each

bypassing for maintenance purposes.

torquemeter indicator presents this information in terms of

percent of engine power being delivered to the main

transmission. Components of the systems include two

pressure chambers, two balancing valves, two pressure

EMER START

transmitters, one high-pressure oil pump, and two dual-

1 ENG 2

needle torquemeter indicators. The system is designed to

measure the oil pressures required to react against the

forward displacement of the main gearbox second-stage

helical gear as a result of the input shaft torque of each

engine. These oil pressures, required to react against the

forward movement of the second-stage helical gear, are

sensed by two pressure transmitters that send electrical

signals to the torquemeter indicators.

Figure 2-8. Engine Emergency Start Switches

2.1.7.9 Engine Starting. During start, as the engine speed

selector is advanced to GRD IDLE, the stopcock opens and

allows fuel to pass through the flow divider and to enter the No.

1 (low pressure) manifold to the nozzles where it is mixed with

compressor-discharge air. As fuel-air mixture leaves the

nozzles, it is ignited by the two igniter plugs in the combustion

chamber and enters a sustained combustion process.

2.1.8 Ignition System. The ignition system consists of a

sealed ignition unit and two igniter plugs on each engine that

provide a spark to ignite the fuel-air mixture. The ignition unit

is on the lower right side of the compressor casing. The ignition

Figure 2-10. Torquemeter

system provides ignition during starting with the ignition switch

at NORM. When gas generator speed increases and the starter

2.1.9.1 Torquemeter Indicators. Two torquemeter

circuit load drops, the automatic dropout relay releases,

indicators (Figure 2-10), one for the pilot and one for the

deenergizing the starter and ignition systems, and combustion is

copilot, are on the instrument panel. Each dual-needle

self-sustained.

indicator marked PERCENT TORQUE contains two

pointers marked

1 and 2 that indicate input torque in

2.1.8.1 Ignition Switches. Two ignition switches, one for

percent of maximum engine power output of each engine.

each engine, on the overhead switch panel (Figure 2-9) are

The electrical pressure torquemeter indicator dials marked

marked IGNITION 1 ENG 2. Each switch has marked

percent torque are marked in units of 5 percent from 0

positions TEST, OFF, and NORM. The switches are normally

percent to 150 percent. The torquemeter indicators operate

2-19

ORIGINAL

NAVAIR 01-230HLH-1

accessory drive. When an overspeed condition occurs, a

The APU fuel system is independent of the helicopter fuel

frequency signal from the No.

1 PMG will actuate a

system, except that it receives its fuel supply from the

solenoid that bleeds compressor discharge pressure (P3)

aircraft’s fuel system. The fuel shutoff valve, a motor-

sensed by the fuel control. With the No.

1 engine in

operated gate valve, is in the right sponson. Normal fuel

accessory drive, two situations will cause activation of this

flow to the APU is accomplished by positioning the APU

ON switch, on the cockpit APU control panel, to ON,

overspeed system. Either No. 1 N_f at approximately 108

which opens the fuel shutoff valve.

percent or N_r at approximately 104 percent or above will

cause the N_f to oscillate. N_r reduction to less than 104

2.1.11.5

(ET) APU Generator System. The APU

percent will restore normal No.

1 engine operation.

generator system, tied into the ac power supply system

(Mechanical engine overspeed protection is not affected by

through the APU generator control unit, supplies power to

AFC 399.)

all components operating off the No.

1 and No.

generators. In addition, to preventing paralleling of the

Note

APU generator with the generator in the basic ac power

system, an APU interlock relay is provided.

With AFC 399, N_r above 104 percent, and

in accessory drive No. 1 engine should sense

2.1.11.6

(ET) APU Operation. Positioning the APU

an overspeed condition and drop to 60- to

CONTROL switch, on the APU control panel, to ON sends

70- percent N_f.

a 28 vdc signal from the battery bus to the ESU. A start

command from the ESU sends 28 vdc to the APU starter

2.1.11

(ET) AUXILIARY POWER UNIT

(APU)

relay and causes the STARTER ON warning light to

SYSTEM. The auxiliary power unit system provides an

illuminate. The ESU also controls ignition and signals the

external source of 115/200 volts, three-phase, 400 Hz, ac

fuel solenoid valve to open. This accelerates the APU to

electrical power in the same manner as an external power

100% speed. When speed reaches 70%, the ESU turns off

unit. The system consists basically of an auxiliary power

the APU starter by transmitting a discrete dc signal closing

unit, an Electronic Sequence Unit (ESU), a control panel,

the APU starter relay. The starter relay drops out and the

an APU fuel system, and a generator system. The APU

STARTER ON warning light goes off. At

95% the

ignition unit is deenergized and after a time delay the APU

system supplies power for instrumentation, minimal

is at 100% speed. The APU ON light will illuminate. The

lighting, and ground operation of the environmental

helicopter’s electrical requirements are supplied when the

control system.

GEN ON switch is positioned ON, and the GEN ON light

illuminates.

2.1.11.1 (ET) Auxiliary Power Unit (APU). The APU,

mounted in a pod attached to the right sponson, supplies

2.2

ENGINE OIL SYSTEM (F0-4)

power to operate the APU generator system. The APU is a

small gas-turbine engine capable of producing

horsepower. It is controlled by the ESU.

Each engine has an independent oil tank and dry sump

full scavenge oil system. Oil is gravity-fed from the tank to

2.1.11.2

(ET) Electronic Sequence Unit (ESU). The

the engine-driven oil pump mounted on the forward right-

ESU is the control element for the APU. The ESU is a

hand side of the engine. The engine-driven pump

microprocessor-based electronic component that monitors

distributes the oil under pressure through a filter to

and controls all APU operations. It monitors speed and

accessory gears and engine bearings. The oil serves both

temperature and outputs signals to control APU functions.

lubricating and cooling purposes and is a completely

The ESU also shuts down the APU when monitored values

automatic system requiring no control action by the pilot.

exceed established operating parameters. Advisory lights

The scavenge side of the pump returns oil through an oil

on the APU control panel indicate system status.

cooler to the oil tank.

2.1.11.3

(ET) APU Control Panel. The control panel

The oil cooler is an oil-to-fuel heat exchanger with an

(Figure 2-5.1), on the cockpit console, contains all of the

associated oil bypass valve. The oil flow through the

controls necessary for APU operation. It consists of an

cooler depends on oil temperature. At lower temperature,

APU ON switch, GEN ON/OFF/RESET switch, GEN ON

the pressure differential across the cooler causes most of

indicator light, FUEL PUMP ON indicator light, and five

the oil to flow through the bypass valve. At higher

warning and indicating lights.

temperatures, the lower viscosity reduces the pressure

differential that closes the bypass valve and causes all of

2.1.11.4

(ET) APU Fuel System. The fuel system

the oil to flow through the cooler. Each engine oil system

supplies fuel to the APU from the helicopter forward fuel

tank. The APU fuel system consists of a APU fuel shutoff

has a 2.7-U.S. gallon capacity tank. The circular tanks are

valve, fuel boost pump, and the necessary tube and hose

around the forward section of each engine. For oil

specification and grade, see Figure 3-7.

lines to convey fuel from the main fuel line to the APU.

2-21

ORIGINAL

NAVAIR 01-230HLH-1

2.3 ROTOR SYSTEMS

stainless steel bonded to the leading edge. The tip caps

have hard nickel bonded to the leading edge.

The rotor systems consist of a single main lifting

rotary wing and an antitorque rotary rudder. Both systems

2.3.1.1 Blade Inspection Method Indicators

are driven by the two engines through the transmission

system and are controlled by the flight controls.

2.3.1 Rotary Wing System. The rotary wing system

WARNING

consists of the rotary wing head assembly and the rotary

wing blades. The rotary wing head assembly, mounted

directly above the main gearbox, consists of a hub

When black is visible in the indicator, it may

assembly and a star assembly. The hub assembly,

be an indication of blade damage that is a

consisting of five sleeve-spindle assemblies and five

flight hazard. The cause of the black

hydraulic dampers clamped between two parallel plates, is

indication shall be determined before flight.

splined to the rotary wing drive shaft. The root ends of the

five rotary wing blades are attached to the sleeve-spindle

A cylindrical BIM indicator (Figure

2-11) is in the

assemblies that permit each blade to flap vertically, hunt

root end plate of each main blade, an air valve is in the

horizontally, and rotate on their spanwise axis to change

backwall of the spar. The pressure indicator has a

the angle of incidence. Antiflapping restrainers limit the

transparent cover through which color indication can be

upward movement of the blades. Droop stops limit the

observed to determine blade serviceability. The indicator

downward position of the blades. Both are in operation

that is compensated for temperature changes compares a

when the blades are stopped or turning at low speed. When

reference pressure built into the indicator with the pressure

the rotor system is accelerated, centrifugal force

in the blades spar. When the pressure in the blade spar is

automatically releases the antiflapping restrainers at about

within the required service limits, indicating the blade is

25 to 30 percent and the droop stops at about 65- to 75-

serviceable, three white stripes show in the indicator. If the

percent rotary wing speed. During rotor deceleration, the

structural integrity of the spar is impaired, nitrogen

droop stops will seat at about 50 to 60 percent and the

pressure will decrease. If the pressure in the blade spar

antiflap restrainers at about 30-percent rotary wing speed.

drops below the minimum permissible service pressure,

The hydraulic dampers lessen hunting movement of the

the indicator will be actuated and will show three black

blades about the vertical hinges as they rotate, prevent

stripes. To check the integrity of the BIM indicator,

shock to the blades when the rotary wing is started or

depress the manual test lever until a black indication

stopped, and aid in the prevention of ground resonance.

appears, and then release the lever.

The angle of incidence (or pitch) of the rotary wing blades

is controlled by the rotary wing flight control system that

2.3.1.2 IBIS

(In-Flight Blade Inspection System).

is connected to the blades through a swashplate assembly

The IBIS system consists of the five main rotor blade spars

below the hub assembly. The swashplate assembly consists

(individually pressurized with nitrogen), a pressure

of an upper (rotating) swashplate that is driven by the

indicator

(with radioactive source) on the root of each

rotary wing hub, and a lower (stationary) swashplate that is

blade spar, a radiation detector on the main rotor fairing

secured by a scissors to the main gearbox. Both

assembly, a signal processor with a test panel in the

swashplates are mounted on a ball-ring and socket

baggage compartment, and an amber blade pressure

assembly that keeps them parallel at all times but allows

caution light located on the caution panel. The nitrogen

them to be tilted, raised, or lowered simultaneously by

with each main rotor blade spar is pressurized to

components of the rotary wing flight control system that

approximately 10 psi at an ambient temperature of 10 to 24

connect to arms on the lower

(stationary) swashplate.

degrees. The associated spar pressure indicator displays

Cyclic or collective pitch changes introduced at the

two white stripes when the pressure is safe. If the spar

stationary swashplate are transmitted to the blades by

pressure drops to below the safe range, two black stripes

linkage on the rotating swashplate. The five all-metal

will display in the spar pressure indicator windows,

rotary wing blades are made of aluminum alloy, with the

indicating that nitrogen is escaping through a crack in the

exception of the forged steel cuffs on the blades that attach

spar or through a faulty seal.

the blades to the sleeve-spindle assemblies on the rotary

wing hub assembly. The main spar of the blade is a

hollow aluminum alloy extrusion that forms the leading

edge of the blade. Individual pockets constructed of sheet

aluminum alloy form the trailing edge of the blades. The

pockets are bonded to the spar. The rotary wing blades

have an abrasion/erosion strip of hard nickel-plated

2-22

ORIGINAL

NAVAIR 01-230HLH-1

to-test pushbutton on the test panel will result in the

following indications:

WARNING

Note

When black is visible in a spar pressure

indicator, hazardous blade damage may

In each bit switch mode, the press-to-test

exist. The unsafe blade shall be removed

pushbutton must be held down for a

from service until the cause of the unsafe

minimum of 5 seconds to allow the signal

indicator is positively found and corrected.

processor to stabilize.

A spar pressure indicator displaying two white (safe)

Selecting DET with the bit switch and pressing

stripes can be checked by pressing the manual test

the press-to-test pushbutton simulates the

pushbutton in the indicator, which causes two black stripes

detectors normal sensing of its own low radiation

(unsafe indication) to be displayed. Once two black stripes

rate and illuminates the green safe light. This is

are displayed, regardless of the cause, the reset button on

an operational checkout of the detector.

the spar pressure indicator must be pressed with the clear

plastic test/shipping cover in place over the spar pressure

Selecting LOW WARN and pressing the

indicator to prevent the escape of radiation.

pushbutton simulates loss of the detectors own

radiation rate and lights the blade pressure caution

The radioactive source is shielded when the spar

light and the red warning light. This is a

pressure indicator displays white

(safe).

When the

functional checkout of the signal processor.

indicator moves to the black

(unsafe) position, the

radioactive source moves out of the shielded position and

Selecting safe and pressing the pushbutton

is exposed. As the rotating unsafe blade then passes over

simulates the detectors normal sensing of its own

the radiation detector, the rays released by the blades’

low radiation rate and illuminates the green safe

radioactive source are sensed by the detector, which sends

light. This is a functional checkout of the signal

an electrical signal to the signal processor, which in turn

processor.

deenergizes a relay in the processor and causes the blade

pressure caution light to illuminate.

Selecting HIGH WARN and pressing the

pushbutton simulates the detector sensing a high

To ensure that the radiation detector is capable of

radiation rate from a spar pressure low indicator

detecting radiation, a small radioactive source incorporated

with low pressure. The blade pressure caution

in the detector continuously emits radiation at a lower rate

light illuminates and the red warning light

than the radioactive source in the spar pressure indicator.

illuminates. This is a functional checkout of the

The radiation detector continuously senses its own

signal processor.

radiation and sends a signal to the signal processor to

energize a relay that extinguishes the BIM caution light.

The No. 2 primary ac and the No. 2 primary dc buses

2.3.2

Bifilar Absorber. The main rotor system has a

furnish power to the system through two circuit breakers,

bifilar

absorber assembly (helicopters modified by AFC

both marked blade pressure, on the pilot circuit breaker

403) to reduce fatigue stress and improve the overall

panel. In addition to damage causing a loss of pressure in

vibration comfort level throughout the helicopter. The

the spar, failure of the signal processor, or loss of 115-vac

bifilar absorber assembly secured to the main rotor hub

power will each result in illumination of the BIM caution

consists of a five-pointed, star-shaped, aluminum forging

light.

with a 17-pound weight attached to each star point.

2.3.1.3 IBIS Test Panel. An IBIS test panel is mounted

2.3.3 Rotary Rudder System. The system consists

on the signal processor, which is located on the left-hand

of the rotary rudder assembly and rotary rudder blades.

side of the cabin wall. The test panel is used for built-in

The rotary rudder assembly mounted at the upper end of

tests (bits) to verify operation of the radiation detector and

the pylon consists of a rotary rudder hub and the pitch-

the signal processor. A bit selector switch, a press-to-test

changing mechanism. The splined hub is supported and

pushdown, a red warning light, a green safe light, and a

driven by the horizontal output shaft of the tail gearbox.

decal that lists the bit procedures and associated light

The five rotary rudder blades are attached to the rotary

indications are mounted on the panel. Pressing the press

rudder hub by flapping hinges and spindles so that they are

2-23

ORIGINAL

NAVAIR 01-230HLH-1

Figure 2-11. BIM Indicator

free to flap and rotate about their spanwise axis for pitch

2.4

TRANSMISSION SYSTEM

variation. The blade pitch-changing mechanism transmits

rotary rudder control pedal movements to the rotary rudder

The transmission system

(Figure

2-12) consists of

blades through the hollow horizontal output shaft of the

three gearboxes that transmit power to the rotary wing and

tail gearbox. The five all-metal rotary rudder blades are

rotary rudder. The three gearboxes are the main gearbox,

constructed of a single aluminum pocket bonded to a C-

intermediate gearbox, and the tail rotor gearbox.

shaped spar. The rotary rudder negative force gradient

system is installed to relieve the pilot of rotary rudder

2.4.1 Main Gearbox.

The MGB

(Figure

2-13),

forces created by aerodynamic loads when the auxiliary

mounted above the cabin an of the engines, interconnects

servo system is inoperative. The system applies a force to

the two engines through a main gearbox to the rotary wing.

cancel the aerodynamic loads only when the rotary rudder

Gearing reduces engine rpm at a ratio of approximately 93

is operating at normal speeds. Because of this, when the

1 for driving the rotary wing. Engine torque is

system is checked on the ground with rotary rudder

transmitted by the main gearbox to the rotary wing drive

stationary and the auxiliary servo off, a negative spring

shaft to drive the rotary wing, and aft to the intermediate

centering effect is created. The tendency of the pedals is

gearbox and then to the rotary rudder gearbox to drive the

then to go normally to either extreme. Under these

rotary rudder. The main gearbox accessory section (Figure

conditions, considerable force is required to push the

2-14) at the rear of the main gearbox lower housing drives

rotary rudder pedals from the extreme positions; however,

the primary, utility, and auxiliary hydraulic pumps, the

the forces will decrease as the positions approach neutral.

high-pressure torquemeter oil pump; and the two

The initial force to move the pedals toward the right from a

generators. Dual oil pumps are installed on the accessory

full left position is between 30 and 40 pounds. With the

section. These pumps increase reliability through better

primary servo ON, auxiliary servo OFF, and the rotary

lubrication and permit flight to be continued if one pump

wing head stationary, pedal motion will cause the col-

fails. A freewheeling unit at each engine input to the main

lective pitch lever to move up for a right pedal motion and

gearbox permits the rotary wing to autorotate without

down for a left pedal motion.

engine drag in case of engine (or engines) failure, or when

engine rpm decreases below the equivalent of rotor rpm.

2-24

ORIGINAL

NAVAIR 01-230HLH-1

2.4.1.1 Accessory Drive Rotor Lockout System.

This system permits the pilot to use engine power to drive

the accessory section

(hydraulic pumps, oil pumps,

generators, etc.) of the main gearbox on the ground

without rotating the rotary wing head. No. 1 engine is

used to drive the accessories without turning the rotary

wing head. A switch allows the pilot to position the linear

actuator and divert power to either the accessory section

or rotary wing.

2.4.1.2 Accessory Drive Panel. This panel (Figure 2-

15) is beside the engine control quadrant. The accessory

drive switch with marked positions FLIGHT and

ACCESS DR is a lever-lock switch that must be pulled

out before it can be moved from one position to the other.

It allows the No. 1 engine to drive the accessory section

of the main gearbox before starting the No. 2 engine and

engaging the rotary wing head and/or when No. 2 engine

rpm is lower than No. 1 engine rpm. Prior to starting the

No.

1 engine, placing the switch to ACCESS DR

positions the rollers in the input freewheel unit, permitting

the No. 1 engine to drive the accessory section of the

main gearbox. Also, at this time, the accessory drive

(rotor lockout) warning light marked ON will go on and

will remain on until the rollers are repositioned to drive

the rotary wings. (On helicopters modified by AFC 401,

the linear actuator will take 6 to 7 seconds to complete

shift. Confirmation is provided by the appropriate light

indicator on the accessory drive panel.) When both

Figure 2-15. Accessory Drive Switch Panel

engines are operating and the rotor is being driven by the

No. 2 engine, No. 1 engine speed should be reduced to

2.4.1.3 Accessory Drive Warning Light

GRD IDLE. When these conditions have been met,

placing the accessory drive switch to FLIGHT energizes

an actuator that repositions the input freewheel unit

rollers, permitting the No. 1 engine to also drive the rotary

wing shaft. An accessory drive limit switch is incorporated

in the ground idle range of the speed selection. This limit

switch prevents switching from ACCESS DR to FLIGHT

On helicopters modified by AFC 401, when

utilizing the override switch to reposition

without first retarding the No. 1 engine speed selector

the actuator, the No.

2 engine must be

lever to the GRD IDLE range. After the rollers are

repositioned, the accessory drive (rotor lockout) warning

driving the rotors at 102- to 104-percent N_r

and the No.

1 speed selector must be

light will go off. (On helicopters modified by AFC 401,

positioned at GRD IDLE to prevent

two interlocks prevent the linear actuator from

repositioning. The No. 2 engine must be driving the rotors

damage to the MGB. Use of this override

switch is not a normal procedure.

at 102- to 104-percent N_r and the No. 1 speed selector at

ground idle. A guarded two-position override switch is

also provided.) The accessory drive switch circuit

Note

operates on dc power and is protected by a circuit breaker

In the event of an input freewheeling unit

marked ACCESS DRIVE on the center circuit breaker

panel (Figures 2-23 and 2-25).

failure while shifting from accessory drive

to flight position on the No. 1 engine (with

a positive indication of successful shift), no

indication of torque on No.

1 eng. N_f

marrying with N_r will be evident after the

shift, regardless of speed selector lever

position.

2-27

ORIGINAL

NAVAIR 01-230HLH-1

This light (Figure

2-15) is on the accessory drive

2.4.4 Intermediate and Tail Gearbox Oil

panel. It will light when the rollers in the freewheel unit

Systems. Both the intermediate and the tail gearboxes

have been positioned to permit the No. 1 engine to drive

are splash-lubricated from individual sump systems.

the accessory section of the main gearbox. When No. 1

Internal spiral channels ensure oil lubrication to all

engine power is diverted to drive the rotary wing shaft,

bearings. An oil filter plug, drain plug, and oil level

the warning light will go off. The accessory drive warning

window are in each gearbox casting. Oil capacity for the

light on is the only positive indication of the input

intermediate gearbox is about 0.2 gallon and for the tail

freewheeling unit in the accessory drive position, and the

gearbox, 0.4 gallon. For oil specification and grade, see

accessory drive warning light off is the only positive

Figure 3-7.

indication of the input freewheeling unit in the flight

position. The press-to-test light operates on dc power and

2.4.5 Chip Detector Caution Lights. These lights

is protected by the circuit breakers marked PWR and

marked MAIN TRANS CHIP, INTMED TRANS CHIP

TEST in the center circuit breaker panel (Figures

2-23

and TAIL TRANS CHIP are on the CAUTION PANEL

and 2-25) under the general heading of WARNING LTS.

(Figure 2-49). The MAIN TRANS CHIP light serves a

single function in that it visually indicates that one or both

2.4.1.4 Tail Takeoff Freewheel Unit Caution Light.

of the magnetic chip detectors in the main gearbox has

The light marked TAIL TAKEOFF is on the caution

picked up and retained metal particles or chips in the oil.

panel (Figure 2-49). The caution light indicates failure of

The INTMED TRANS CHIP and TAIL TRANS CHIP

the tail takeoff freewheel unit in the main gearbox

lights, when on, indicate that the gearboxes have

accessory drive train. When this failure occurs, the

overheated or the magnetic chip detectors have picked up

accessory section is being driven by the No. 1 engine

and held chips or metallic particles in the oil. The

through-shaft at reduced rpm. No. 1 generator output is

presence of either of these conditions could cause ex-

normally used to sense this reduction in rpm that results in

cessive wear and/or premature failure of the gearboxes.

a reduction in generator frequency. In case the No. 1

The system operates on dc power and is protected by a

generator has failed, the No. 2 generator output is used to

circuit breaker marked CHIP DET on the center circuit

sense this rpm reduction. When the generator frequency is

breaker panel.

reduced to less than that produced through the freewheel

unit drive, the caution light will go on. Failure of the No.

2.4.6 Transmission Oil Systems

1 engine subsequent to the caution light going on would

result in loss of the equipment driven by the accessory

2.4.6.1 Main Gearbox Oil System (FO-5). Primary

section. The system operates on dc power. The sensing

and secondary oil pumps are for lubrication. The primary

unit is in the electronics compartment.

oil pump is mounted on a common shaft between the

torque indicating system oil pump and the transmission

rear cover. The secondary oil pump is mounted on a

Note

common shaft between the utility hydraulic pump and the

transmission rear cover

(Figure

2-14). Oil is pumped

Flight regimes above 100-percent N_r may

from the gearbox sump through a hose to an oil cooler

preclude detection of an actual tail takeoff

behind the main gearbox. If oil from the sump is less than

freewheel unit failure.

70^o C, the thermostatic bypass valve in the oil cooler

opens and the oil bypasses the radiator and goes directly

2.4.2 Intermediate Gearbox. This gearbox

(see

to the jets. If the oil temperature is over 70^oC, the valve

Figure

2-12) at the base of the rotary rudder pylon

closes, directing oil through the oil cooler to be cooled

contains a bevel gear direct-drive system to change the

before going to the jets. Cooling air enters the forward

direction of the shafting that transmits engine torque to

end of the main gearbox fairing through a screened main

the tail gearbox. The intermediate gearbox is splash-

gearbox cooling air intake

(Figure

2-2) and is forced

lubricated. Screened intermediate gearbox cooling air

through the oil cooler by a blower driven by belts from

inlets (Figure 2-2) in the pylon fairing permit the gearbox

the tail drive shaft. The air is then exhausted through a

to be cooled by the rotary wing downwash.

screened transmission accessories cooling air outlet at the

rear of the fairing. After passing through the oil cooler;

2.4.3 Tail Gearbox. This gearbox (see Figure 2-12)

the oil returns to the main gearbox where it is sprayed

at the upper end of the rotary rudder pylon contains a

onto the gears and bearings through jets built into the

bevel gear reduction-drive system to transmit engine

gearbox castings. An oil filler, reached from the left side

torque to the rotary rudder. The tail gearbox also contains

of the rotary wing fairing, is on the left side of the

part of the pitch change linkage that extends through the

gearbox. A window in the gearbox below the oil filler

hollow horizontal output shaft to the rotary rudder hub.

provides a sight check for the oil level in the main

The tail gearbox is splash lubricated.

gearbox. Oil tank capacity is

16.6 gallons; normal

servicing is 12.6 gallons.

2-28

ORIGINAL

NAVAIR 01-230HLH-1

2.4.6.2 Main Gearbox Oil Pressure Indicator and

2.4.6.4 Main Gearbox Oil Temperature Indicator

Caution Light. The main gearbox oil pressure indicator

and Caution Light. The main gearbox oil temperature

(Figure 2-7) is on the instrument panel. The indicator is

indicator (Figure

2-7) marked XMSN OIL TEMP on the

marked in pounds per square inch, and is activated by a

instrument panel is graduated in degrees Celsius. The

pressure transmitter connected to the gearbox inlet

indicator is connected by direct current from the primary

lubrication line. The main gearbox oil pressure indicator

bus to an oil temperature bulb next to the main gearbox

operates on 26 vac and is protected by a circuit breaker

oil outlet port and is protected by a circuit breaker marked

marked XMSN OIL PRESS. The main gearbox oil low-

XMSN OIL TEMP. The main gearbox oil temperature

pressure caution light marked TRANS OIL PRESS is on

caution light is on the caution panel and is activated by a

the CAUTION PANEL (Figure 2-49) and is actuated by a

temperature sensor on the inlet lubrication line. The

pressure switch in the forward part of the main gearbox.

amber caution light operates on dc power and is protected

The amber caution light operates on dc power and is

by a circuit breaker marked WARNING LTS PWR on the

protected by a circuit breaker marked WARNING LTS

center circuit breaker panel

(Figures

2-23). The

PWR on the center circuit breaker panel. The light will

transmission oil temperature caution light will go on when

go on when the main gearbox oil pressure drops below 3-

the transmission oil temperature at the oil cooler exit is

1/2 ±1 psi at the pressure switch at the forward right

over 120^o C. The different locations of the temperature

corner of the main gearbox at the point furthest from the

sensors for the gauge and light allow the pilot to monitor

pressure pumps. The different locations of the pressure

the gearbox operation by means of the gauge and the oil

sensors for the gauge and caution lights were incorporated

cooler operation by means of the caution light. Thus, if a

to warn the pilot of an oil blockage within the gearbox

malfunction occurs in the oil cooler (blockage, fan belt

that may not be indicated on the pressure gauge.

failure, etc.), the caution light will go on before the

gearbox oil temperature rises to a dangerous level. The

2.4.6.3 Main Gearbox Emergency Lubrication

120^o C for the caution light limitation is sensed after the

System. The ELS will permit continued main gearbox

oil cooler, and the

145^oC limitation noted on the oil

operation for a limited time following failure of the main

temperature indicator is sensed before the oil cooler.

lubrication system. The ELS was designed to allow

continued flight for up to 30 minutes following a failure

2.4.7 Rotor Brake. A hydraulically activated rotor brake

of the main lubrication system. In the event of a sudden

mounted on a brake shaft forward of the main gearbox

massive loss of lubricating oil, vibrations and/or other

stops the rotation of the rotor system and prevents its

secondary indications that require immediate action may

rotation when the helicopter is parked. The rotor brake

occur within 2 minutes. The exact duration of safe flight

consists of a hydraulic cylinder and lever pressure gauge

using ELS will vary depending upon various conditions,

hydraulic brake cylinders, and a brake disc. The rotor

including but not limited to, mode of main lube failure,

brake hydraulic cylinder and lever on the pilot

aircraft weight, aircraft pitch attitude, balanced flight

compartment ceiling operate independently from the

condition, and power requirements. An enlarged sump at

hydraulic systems. A spring-loaded accumulator

the base of the main gearbox provides an additional 1.6

connected to the rotor brake hydraulic lines at the forward

gallons of oil for the ELS. Upon a malfunction of the

end of the transmission compartment assures continuous

main lubrication system, a pressure-sensitive valve will

hydraulic pressure after the rotor brake lever is applied.

activate when the oil pressure at the input sleeve bearings

The rotor brake hydraulic cylinder is gravity-fed with

in the high-speed section of the main gearbox falls below

hydraulic fluid from the utility hydraulic system reservoir.

12 to 15 psi that corresponds to 20 to 25 psi on the oil

In case of a broken or leaking hydraulic line from the

pressure gauge. This valve directs the ELS oil to the

utility hydraulic system reservoir, the rotor brake

sleeve bearings as well as the torque system. ELS oil

hydraulic cylinder contains enough fluid for braking the

pressure is provided by the torquemeter pump at the rate

rotary wing system. The hydraulic brake cylinder is on

of 0.7 gpm. Although the pressure -sensitive check valve

supports attached to the main gearbox. The brake disc

prevents oil from the torquemeter pump from being

positioned on the input shaft of the main gearbox has a

supplied to other parts of the transmission, the ELS is not

toothed edge that is engaged by the blade positioner drive

a closed-loop system. ELS system oil returning from the

unit to turn the rotary wing in the blade positioning cycle.

input shaft bearing passes through the main sump on its

way to the ELS sump. When a leak occurs in the main

lube oil system, the primary and secondary lube oil

pumps, if operating, will deplete a portion of the

remaining ELS oil as it passes through the main sump.

While the ELS operates, oil pressure, oil temperature, and

the torque indicator will not register accurately. A low

but steady torque reading rather than zero torque reading

will confirm that the ELS is operating.

2-29

ORIGINAL

NAVAIR 01-230HLH-1

2.4.7.2 Rotor Brake Pressure Gauge. The hy-

draulically actuated gauge is to the rear of the rotor brake

lever (Figure 2-16) on the pilot compartment ceiling. The

WARNING

reading that is indicated by the needle indicates psi X 100.

A decal, marked ROTOR BRAKE PRESS.,

ACTUATING RANGE 350-500 PSI., ENGINE START

The rotor brake will not prevent rotor movement

320 PSI. MIN., PARKED POS., and RANGE 250/600

with the No.

1 engine in flight position above

PSI., is next to the rotor brake pressure gauge. A pressure

ground idle or with the No. 2 engine above ground

of 320 psi or more is needed before an engine start can be

idle. Personnel injury and/or helicopter damage

initiated with the blades folded.

may occur as a result of inadvertent rotor

engagement.

2.4.7.3 Rotor Brake Caution Light. This light on the

CAUTION PANEL indicates that either the manual or

2.4.7.1 Rotor Brake Cylinder and Lever. A rotor

automatic rotor brake is on; however, the automatic rotor

brake lever (Figure 2-16) is connected directly to the rotor

brake is not applied until the blade fold master switch is

brake hydraulic cylinder to the right and forward of the

activated. The light will go off when the rotor brake is

disengaged.

overhead switch panel on the pilot compartment ceiling.

The rotor brake is applied by pulling down and forward as

indicated on the decal aft of the lever on the upper

2.5

FUEL SYSTEM

structure. The decal is marked TO ENGAGE ROTOR

BRAKE PUSH LEVER FORWARD. The decal also has

2.5.1 Internal Fuel System. These helicopters are

an arrow pointing forward. When actuated, a lock-lever at

equipped with two pressure-type fuel systems (FO-7). The

forward tank system supplies fuel to the No. 1 engine and

the forward outboard side of the cylinder locks the brake

lever in the applied (forward) position. To release the

the aft tank system supplies fuel to the No. 2 engine. On

rotor brake, reposition the lock-lever aft and swing the

the ET the forward tank also supplies fuel to the auxiliary

rotor brake lever aft and up against the bottom of the

power unit (APU). The forward and aft systems each

cylinder until it snaps into place. The lockpin may be

consist of a tank with two bladder type cells, a collector

made inoperative by turning it until it remains in the OUT

can equipped with two boost pumps, an ejector system

that continually fills the collector can with a boost pump

position. For normal shutdown, the rotor brake should be

applied firmly and smoothly. As rotor rpm approaches

on, a main line filter, and a firewall shutoff valve. The

zero, rotor brake pressure should be reduced in order to

collector can boost pump and ejector arrangement provide

ease rotor blades to a stop, precluding any tendency of

for a minimum of unusable fuel. A crossfeed line between

whip stopping. When the rotary wing blades are folded,

the two pressure-type systems permits fuel from both the

the rotor brake lever must be on to open interlock for a

forward and aft tanks to be directed to one engine during

single-engine operation, or fuel from one tank to supply

normal start. For emergency shutdown, the rotor brake

lever may be forced forward into the full ON position

both engines.

after closing the engine speed selectors. In case of an

emergency, the rotor brake, when fully applied, is

The center tank replaces consumed fuel in the forward

designed to stop the rotors from

77-percent N_r in 14

and aft tanks. The center tank consists of a single

seconds with engines at idle and from 91 percent in 20

bladder-type cell, two ejectors, and a common opening

between the forward and center tanks. The opening

seconds with engines at idle.

allows fuel to spill freely into either tank. One center tank

ejector is associated with the forward tank and the other

with the aft tank. The center tank ejector associated with

the forward tank will pump fuel into the forward tank

when one or two forward tank boost pumps are on. This

is the only means of transferring fuel from the center to

forward tank once the fuel level drops below the common

opening level. The center tank ejector associated with the

aft tank will pump fuel into the aft tank when one or two

aft tank boost pumps are on and the float valve in the aft

tank is open. The aft tank float valve opens when the aft

tank fuel level drops below 600 to 900 pounds.

Figure 2-16. Rotor Brake Lever

When the aft tank fuel system is operating and the aft

tank float valve is closed, fuel flow from the aft tank

system to the center tank aft ejector transfers fuel from the

aft tank into the center tank. To prevent overfilling the

2-30

ORIGINAL

NAVAIR 01-230HLH-1

center and forward tanks, a float shutoff valve is

Placing the switch in the OPEN position electrically

incorporated in the center tank aft ejector line. The fuel

opens the crossfeed valve that connects the systems. The

management panel (Figure 2-17) on the instrument panel

crossfeed system may be used to supply fuel under

contains the four fuel boost pump switches, the boost

pressure from both tanks to any one or both engines. The

pump failure warning lights, the crossfeed switch, and the

crossfeed system does not transfer fuel between tanks.

two firewall shutoff switches. The tanks may be filled by

either a pressure-refueling system or the conventional

2.5.1.3 Fuel Boost Pumps. Two boost pumps in each

gravity feed through the filler necks. Each tank has a fuel

collector can in the fuel tanks supply fuel under pressure

quantity gauge that indicates the quantity of fuel in

to the two independent fuel systems. Normally, fuel flows

pounds. See Figure 2-22 for fuel quantity data and Figure

from the tanks through a fuel filter, check valve, and

3-7 for fuel specification and grade.

firewall shutoff valve to the engine driven fuel pump and

then to the engine fuel control unit. Should to engine-

2.5.1.1 Fuel Shutoff Valve Switches. Two fuel

driven fuel pump fail, the engine will fail because of fuel

shutoff valve switches marked FIREWALL VALVE with

starvation.

the marked positions OPEN and CLOSE are on the fuel

management panel. These switches control the fuel

2.5.1.4 Fuel Boost Pump Operation

shutoff valves overhead in the cabin in front of the engine

compartment. Placing the switches in the CLOSE

Note

position shuts off the flow of fuel to the engines. In case

of electrical failure, the valves will remain in the last

When starting No. 1 engine with an exter-

energized position. The switches and valves are powered

nal dc power source or the battery, the

by the primary bus, and protected by circuit breakers on

boost pumps will be inoperative because of

the center circuit breaker panel marked 1 ENG 2 under

their ac power source requirement. If all

the general headings FUEL SYSTEM and VALVES.

fuel lines to the engine are full, the engine-

driven fuel pumps will be capable of

supplying sufficient fuel for starting.

2.5.1.5 Fuel Boost Pump Switches. Four boost

pump switches on the fuel management panel control the

fuel flow to the engine. The switches are in sets of two;

those for the forward tank are marked FWD TANK while

the two for the aft tank are marked AFT TANK. Above

each switch is a number 1 or 2 to designate the pump in

the tank controlled by that switch. Each switch has two

marked positions: PUMP (ON position) and OFF. The

boost pumps in the collector can in the forward cell of

each fuel tank supply fuel to the engine-driven pump

when the switches are in the PUMP position

(FWD

TANK and AFT TANK). The pump switches are

controlled by dc power and are protected by four circuit

breakers on the center circuit breaker panel. The boost

pumps are operated on ac power with each pump pro-

tected by a circuit breaker.

2.5.1.6 Fuel Boost Pump Failure Warning Lights.

Four pressure switches, two installed in each tank, are

Figure 2-17. Fuel Management and Fuel Quantity

connected to the pressure feed line from each boost pump.

Panel

The pressure switches are connected by direct current

through the warning light circuit breaker on the center

2.5.1.2 Crossfeed Switch. A fuel crossfeed switch

circuit breaker panel and are marked PWR and TEST

marked CROSSFEED is on the fuel management panel.

under the general heading WARNING LTS. When the

The switch has marked positions CLOSE and OPEN and

boost pumps are first turned on or the boost pump

actuates a valve that operates on dc power and is

switches are tested, the fuel pump failure warning lights

protected by a circuit breaker marked X FEED on the

will go on and then off. They are lit until pressure is built

center circuit breaker panel under the general headings

up in the system. The switches close if the pumps fail, and

FUEL SYSTEM and VALVES. Normally, the switch is

the warning lights marked PUMP FAILURE on the fuel

in the CLOSE position. With the switch in the CLOSE

management panel will light. Pressure must decrease to

position, the No. 1 engine receives fuel from the forward

approximately 16-1/2 psi to energize the warning circuit.

tank and the No. 2 engine receives fuel from the aft tank.

2-31

ORIGINAL

NAVAIR 01-230HLH-1

2.5.1.9 Fuel Filter Bypass Caution Lights. The fuel

WARNING

filter bypass caution lights will light whenever fuel bypass

is imminent. This normally occurs when a pressure

differential of approximately 1.1 to 1.7 psi is sensed at the

filter. As the contamination increases and the pressure

If a fuel boost pump failure warning light

differential increases to 1.9 to 2.3 psi, the bypass valve

goes on, activate the remaining boost pump

opens and, at this point, the filter is being bypassed. The

for that tank before securing the affected

caution lights are marked FWD FUEL BYPASS and AFT

pump. If the fuel boost pump failure

FUEL BYPASS, indicating which fuel system is in a

warning light goes off when both pumps

condition of impending filter bypass. The lights operate

are activated, there may be a fuel leak.

on dc power and are protected by the WARNING LTS

PWR circuit breaker on the center circuit breaker panel.

2.5.1.7 Fuel Low-Level Caution Lights

2.5.1.10 Pressure Fueling-Defueling System. The

system is a single-point fueling-defueling system. The

pressure-refueling filler cap (Figure

2-18) is on the right

side of the fuselage, inside the step below the cargo door.

The filler cap is marked CAP-FUEL FILLER SINGLE

POINT SERVICNG. To fill the fuel tanks, the refueling

When the fuel low-level caution lights go

nozzle is connected to the adapter connection and fuel is

on, attitudes of over

6^o noseup should be

pumped through the fuel lines and the fueling and

avoided because of the possibility of fuel

defueling valves in the forward and aft fuel tanks. The

starvation.

fueling and defueling valves are normally closed; for

refueling, the valves open with pressure at the inlet until

The low-level caution lights are on the CAUTION

high-level shutoffs close the valves. The high-level

PANEL on the instrument panel. The caution lights,

shutoff valve for the forward tank is in the center tank.

operating on dc power, are tested by the master TEST

This allows both the forward and center tanks to fill

button on the caution panel and are protected by circuit

simultaneously through the common opening. When the

breakers marked LOW LEVEL FWD and AFT on the

center tank is full, the refueling process is stopped for

center circuit breaker panel.

those tanks. For defueling, the valves open with vacuum

at the inlet until the low-level shutoffs close the valves.

On UH-3H helicopters, the FWD FUEL LOW and

The low-level shutoffs are only in the forward and aft

AFT FUEL LOW caution lights for the forward and aft

tanks.

tanks will go on when approximately 210 to 280 pounds

per tank remain in a

3^o nosedown attitude, or between

2.5.1.11 Pressure-Refueling

Switches.

The

170 and 200 pounds per tank remain when in a hover.

switches marked PRI TEST and SEC TEST are on a panel

(Figure

2-18) marked PRESS REFUELING PRECHECK

2.5.1.8 Fuel Quantity Gauges and Test Switch.

next to the refueling-defueling adapter connection. The

The fuel quantity gauges on the instrument panel above

switches are used to check the reliability of the fuel high-

the fuel management panel indicate the fuel quantity in

level shutoffs. The panel contains information on the

each tank in pounds. UH-3 helicopters have three quantity

refueling precheck that must be followed prior to using

gauges for the forward, center, and aft tanks, respectively.

the pressure refueling system.

The tank unit capacitance system used in this helicopter is

practically unresponsive to volumetric changes resulting

from various temperatures. The dielectric between the two

electrodes will vary as the fuel varies. The fuel quantity

gauges are calibrated to measure this voltage differential

in pounds of fuel. Fuel quantities are shown in the fuel

quantity data tables. The fuel quantity indicating system

may be tested by pressing the fuel quantity gauge test

switch marked FUEL GAGE TEST to the left of the fuel

management panel. Pressing the button-type switch will

induce a current reversal that causes the needles to turn to

zero. Upon release of the test switch, the normal current

should cause the needles to return to the previous reading.

This test shows that the fuel quantity indicating system is

operating correctly. The fuel quantity indicating system

Figure 2-18. Single-Point Servicing Panel

operates on ac power and is protected by circuit breakers.

2-32

ORIGINAL

NAVAIR 01-230HLH-1

2.5.2

(NON ET) Helicopter In-Flight Refueling

level at which the aft tank float valve opens (600 to 900

System

(HIFR). Helicopters are equipped with a

pounds), the center tank aft ejector starts to transfer fuel

system for refueling from a surface vessel while in flight.

into the aft tank.

Fuel is pumped on board the helicopter under pressure

through a Wiggins quick-disconnect nipple. The Wiggins

For operation under unusual conditions, the fuel

quick-disconnect nipple is in the cabin floor near the

quantity in the tanks can be equalized by operating both

sliding cabin door. The nipple is provided with a cover

engines from either the forward or aft tank systems. This

marked REPLACE COVER AFTER FILLING.

can be done by opening the crossfeed and turning on both

separate grounding receptacle is provided at this station.

boost pumps in the tank system being used, and one boost

A filter is incorporated in the system to prevent

pump on in the tank system not being used. The tank

contaminated fuel from entering the helicopter. In

with two boost pumps on will supply fuel to both engines,

addition, the station is equipped with a dome light with a

because of the greater pressure furnished by the two

control switch. See Chapter 9 for Helicopter In-Flight

pumps that close the check valve downstream of the tank

Refueling Procedures.

with only one pump operating.

2.5.3 Fuel System Management. For normal

2.5.4 Fuel Dump System. This system will dump fuel

operation of UH-3 helicopters (Figure

2-19) with cross-

from a single tank at a rate of 400 to 500 pounds per

feed closed and one boost pump on in each tank, the

minute, and simultaneously from both tanks at approxi-

forward and center tank will furnish fuel to the No. 1

mately 800 pounds per minute. Fuel will stop dumping in

engine and the aft tank will furnish fuel to the No. 2

either tank when the associated tank's fuel low caution

engine. This causes the aft tank quantity gauge to drop

light goes on, since the caution light sensor and the dump

more rapidly than the forward and center tank gauges. In

standpipe are at approximately the same level. The tank's

addition, as long as the center tank float shutoff valve is

dump pump should be turned off at this time to prevent

open and the aft tank float valve is closed, fuel will be

the pump from overheating. If the pump is not turned off,

transferred at a low rate from the aft tank to the center and

a thermal device in the pump will automatically turn it off

forward tanks. This also causes the aft tank quantity

and the pump will have to be removed to reset the thermal

gauge to drop more rapidly than the center and forward

device.

tank gauges. Once the aft tank fuel is depleted below the

Figure 2-19. Fuel System Management

2-33

ORIGINAL

NAVAIR 01-230HLH-1

The system

(FO-7) uses the center point

2.6 ELECTRICAL POWER SUPPLY SYSTEM

pressure-fueling plumbing to dump fuel overboard

(FO-8)

through a tube near the right side of the tailwheel. The

forward and aft tanks each have a dump pump and dump

Two ac generators supply power to the electrical

valve that are off and closed respectively when not

system. Four transformers provide 26-volt single-phase ac

dumping. A system dump valve incorporated in the center

power. Two rectifiers provide

28-vdc control and

point fueling system is closed when not dumping to

operating power. One inverter supplies 115-vac power.

prevent dumping when pressure fueling. When dumping,

One battery supplies 24-vdc power.

the high-level shutoff sensor is actuated and closes the

fueling/defueling valve to prevent dump fuel from

2.6.1 Alternating Current Power Supply System.

reentering the tank.

Ac power is supplied by two generators designated as

Nos.

1 and

2. Associated system components are

The primary dc bus furnishes control power to the

designated in a similar manner. System operation is

tank dump systems through two FUEL SYSTEM circuit

automatic; control switches on the overhead switch panel

breakers on the center circuit breaker panel marked FWD

and monitoring light capsules on the caution and advisory

and AFT above the general heading DUMP. The forward

panels are provided. Normally, associated primary bus

tank dump pump operating power is furnished by the No.

loads are assumed by each generator and the monitored

1 primary ac bus through a circuit breaker on the copilot

bus load is powered by the No. 2 generator. Primary bus

circuit breaker panel marked DUMP FWD under the

loads are those that are essential for night and/or

general heading FUEL PUMP. The aft tank dump pump

instrument flight and for operation of equipment

operating power is furnished by the No. 2 primary ac bus

necessary for mission requirements. The monitor bus load

through a circuit breaker on the pilot circuit breaker panel

is not essential for this type of flight. If either generator

marked DUMP AFT under the general heading FUEL

fails, its primary bus load is automatically transferred to

PUMP. The system dump valve is controlled by primary

the remaining generator. With a failed generator, the

dc bus power through a FUEL SYSTEM circuit breaker

monitor bus load is automatically dropped. An external ac

on the center circuit breaker panel marked FUEL SYSEM

power receptacle permits use of an external ac power unit

above the general heading DUMP.

for ground power application.

2.5.4.1 FUEL DUMP Control Panel. The panel

2.6.1.1 Generators. Two

115-/200-volt, three-phase,

marked FUEL DUMP (Figure 2-20) is on the cockpit

400-Hz, brushless PMGs are mounted on and driven by

console. There are two guarded switches on the panel, one

the accessory section of the main gearbox. Generator

with marked positions FWD and OFF and the other AFT

output varies with temperature and altitude. Generator ac

and OFF. Placing the forward tank switch to FWD opens

voltage is delivered to the associated supervisory panel

the tank dump valve, turns the dump pump on, actuates

and generator contactor relay. The permanent magnet

the high-level shutoff to close the fueling/defueling valve,

sections of the generators develop ac power that is

and opens the system dump valve. The aft tank switch

rectified to dc power in the supervisory panel to be used

operates the same way.

in the control circuits.

2.6.1.1.1 (ET) APU Generator. One 115-/200-volt, 3-

phase, 400-Hz generator is mounted on the APU and is

used for instrumentation, minimal lighting, and ground

WARNING

operation of the air conditioning system. The generator

output varies

with temperature and altitude.

Electrical malfunction in either the forward

(Approximately 32 kVA at sea level.)

or aft dump valves or both may cause

inadvertent fuel dumping without pilot

2.6.1.2 Supervisory Panels. The supervisory panels

knowledge. If this happens, rapid fuel loss

designated Nos. 1 and 2 provide control for relays in the

will result.

electrical system. When the No. 1 generator is developing

normal ac power and the generator switch is placed ON,

PMG ac power from the same generator rectified to dc

power by its associated supervisory panel will be used by

the supervisory panel to close the generator contactor

relay. Closing the No. 1 generator contactor relay permits

the No. 1 generator to power the No. 1 primary ac bus and

the supervisory panel to deliver 28 vdc to the ac monitor

bus relay. In addition, it opens the No. 1 generator caution

light circuit causing the light to go off. The No.

Figure 2-20. Fuel Dump Panel

supervisory panel operates the same way to power the No.

2-34

ORIGINAL

NAVAIR 01-230HLH-1

2 primary ac bus and to turn off the No. 2 caution light.

range. If any of the monitored conditions are not normal,

Dc power from the No. 2 supervisory panel also closes

the generator contactor relay will open, taking the

the ac monitor bus control relay that permits 28 vdc from

associated generator off the line, deenergizing the ac and

the No. 1 supervisory panel to close the dc monitor bus

dc monitor buses, and lighting the associated generator

relay and the ac monitor bus contactor relay. Therefore,

caution light. In case of a generator failure, primary ac

28-vdc power is required from both the Nos. 1 and 2

bus loads will be switched automatically to the remaining

supervisory panels to energize the ac and dc monitor

generator.

buses. If either generator fails to produce PMG ac power,

the primary dc bus supplies backup dc voltage to each

2.6.1.3 Generator Switches. These switches are on

supervisory panel through circuit breakers on the center

the overhead switch panel (Figure 2-9) under the general

circuit breaker panel marked 1 and 2 under the general

heading 1 GEN 2 and have the marked positions ON,

heading PMG BACK-UP. The supervisory panels provide

OFF-RESET, and TEST. Placing the switch ON puts the

protection for the electrical system. Three-phase AC

respective generator on the line by closing the generator

power delivered to the panel from its associated generator

contactor relay. The OFF-RESET position turns the

is monitored by the panel at all times for open phase,

generator off and resets the cycle. When the generators

overvoltage, and undervoltage. The panel monitors for

drop off the line because of an overvoltage or

underfrequency (Figure

2-21) when the helicopter is on

undervoltage they will have to be reset; however, when

the ground with its main landing gear struts compressed.

the generators drop off the line because of an

In flight, the underfrequency protection is eliminated by

underfrequency, they will come back on automatically.

action of the microswitch attached to the scissors of each

The TEST position is used for maintenance.

landing gear. In this condition, the generators will remain

on the line throughout the entire normal rotary wing speed

FLIGHT POSITION

ACCESSORY DRIVE POSITION

PERCENT

GENERATOR

PERCENT

GENERATOR

Nf or Nr

FREQ

380.0

100

400.5

104

400.0

105

420.0

109

420.0

FLIGHT POSITION

NOTES

Under frequency protection is locked out during flight.

Generator should pickup (come on the line) between 92.1% and 96.8% N_f/N_r and should drop out

within minus 2% of the pickup value. (These figures assume a 2% tachometer system accuracy.)

ACCESSORY DRIVE POSITION

NOTES

During ground operation frequency is set at 377-383 Hz for pickup, 374-380 Hz for dropout.

Generator should pick up (come on the line) between 95.8% and 101.3% N_f and should dropout

within minus 2% of the pickup value. (These figures assume a 2% tachometer system accuracy.)

Figure 2-21. Generator Frequency at Various Rotor and Power Turbine Speeds

2-35

ORIGINAL

NAVAIR 01-230HLH-1

(JP-4 AT 15.6^oC)

GRAVITY REFUELING

FORWARD TANK

CENTER TANK

AFT TANK

TOTAL FUEL

GALLONS POUNDS GALLONS POUNDS GALLONS POUNDS GALLONS POUNDS

UNUSABLE

2.65

17.2

.83

5.4

2.65

17.2

6.13

39.8

USABLE

344.35

2238.3

147.17

956.6

350.35

2277.3

841.87

5472.2

FULLY

347.00

2255.5

148.00

962.0

353.00

2294.5

848.00

5512.0

SERVICED

PRESSURE REFUELING

UNUSABLE

2.65

17.2

.83

5.4

2.65

17.2

6.13

39.8

USABLE

338.35

2199.3

145.17

943.6

341.35

2218.8

824.87

5316.7

FULLY

341.00

2216.5

146.00

949.0

344.00

2236.0

831.00

5401.5

SERVICED

1. Data basis 0^o fuselage attitude.

2. Fuel density JP-4 = 6.5 pounds per gallon.

(JP-5 AT 15.6^oC)

GRAVITY REFUELING

FORWARD TANK

CENTER TANK

AFT TANK

TOTAL FUEL

GALLONS POUNDS GALLONS POUNDS GALLONS POUNDS GALLONS POUNDS

UNUSABLE

2.65

18.1

.83

5.6

2.65

18.1

6.13

41.8

USABLE

344.35

2341.6

147.17

1000.8

350.35

2382.4

841.87

5724.8

FULLY

347.00

2359.7

148.00

1006.4

353.00

2400.5

848.00

5766.6

SERVICED

PRESSURE REFUELING

UNUSABLE

2.65

8.1

.83

5.6

2.65

18.1

6.13

41.8

USABLE

338.35

2300.7

145.17

987.2

341.35

2311.1

824.87

5599.0

FULLY

341.00

2318.8

146.00

992.8

344.00

2339.2

831.00

5640.8

SERVICED

Data basis 0^o fuselage attitude.

Fuel density JP-5 = 6.8 pounds per gallon.

Figure 2-22. Fuel Quantity Data

2-36

ORIGINAL

NAVAIR 01-230HLH-1

2.6.1.4 Generator Caution Lights. Two generator

2.6.1.8 External Ac Power Receptacle. The

caution lights marked

#1 GENERATOR and #2 GEN-

external ac power receptacle is mounted on the left aft

ERATOR, respectively, are on the caution panel. These

side of the fuselage. This receptacle is used to introduce

lights will go on whenever the associated generator is

115-/200-volt, three-phase,

400-Hz ac power into the

taken off the line by the opening of the generator

helicopter electrical system.

contactor relay that causes the caution light circuit to be

completed. The generator caution lights are powered by

Note

the primary dc bus and protected by circuit breakers Nos.

1 and 2 under the general headings GENERATOR and

On the UH-3H Executive Transport

WARNING LTS on the center circuit breaker panel.

helicopters, the BRIGHT/DIM switch

on the CABIN LIGHTS panel should

2.6.1.5 Inverter. A

100-volt ampere, 115-volt inverter

remain in the DIM position while

is incorporated in the electrical system to provide ac

operating on external power.

power to the fire detection system, fuel quantity

indicating system, electrical autotransformer, and the

An external power monitor panel monitors external

isolation transformer. The inverter is automatically turned

power voltage level, frequency, and phase rotation. If

on when ac power is not present in the helicopter, and the

these are correct, dc power will pass through the external

primary dc bus is activated by an external dc power unit,

power monitor panel and energize the external power

or the battery switch is turned on. Primary dc bus

contactor relay, introducing external three-phase power

operating power for the inverter is protected by a circuit

into the electrical system. A switch marked RESET-ON-

breaker marked INV on the center circuit breaker panel.

OFF under the heading EXT PWR on the overhead switch

Inverter ac power is protected by two circuit breakers on

panel (Figure

2-9) must be ON to allow power to pass

the copilot circuit breaker panel marked PWR A and

through the power monitor panel. At OFF, ac power is

PWR C under the general heading INVERTER.

removed by the power monitor panel. If ac power voltage

level, frequency, and phase rotation are not correct, 28-

2.6.1.6 Transformers. Four transformers in the ac

vdc power will not energize the external power contactor.

system convert 115-volt power from the primary ac buses

When the fault in the external ac power source is cor-

to 26 volts. The Nos. 1 and 2 radio autotransformers are

rected, the EXT PWR switch must be momentarily placed

to RESET to reactivate the external power monitor panel.

autotransformer supplies power to the TACNAV or

Dc control power is protected by a circuit breaker on the

navigation system, LF/ADF, BDHIs, TCDI, and GSDA.

center circuit breaker panel marked EXT PWR. The EXT

PWR ON advisory light will go on when the ac power

The No. 2 radio autotransformer supplies power to the

receptacle door is open.

tacan, UHF-DF, and compass system. The radio

autotransformers are protected by two circuit breakers on

2.6.1.9 External Power Advisory Light. This light

the copilot circuit breaker panel under the general heading

on the advisory panel marked EXT PWR ON will go on

AUTO XMFR and are marked 1 and 2, respectively. The

when the ac external power door is open or when external

electrical autotransformer is powered by the No.

dc power is being supplied to the helicopter.

primary ac bus or the inverter and supplies power to the

No.

1 engine oil pressure indicator, primary servo

2.6.1.10 Ac Circuit Breakers. Ac circuit breakers are

hydraulic pressure indicator, utility hydraulic pressure

on the pilot and copilot circuit breaker panels (Figure 2-

indicator, and the No. 1 engine torquemeter. The elec-

23).

trical autotransformer is protected by a circuit breaker on

the copilot circuit breaker panel marked AUTO XMFR 0

2.6.2 Direct Current Power Supply System.

C. The isolation transformer is powered by the No. 2

Dc power is supplied by two rectifiers, designated as Nos.

primary ac bus or the inverter and supplies power to the

1 and 2 that are powered by the Nos. 1 and 2 primary ac

No. 2 engine oil pressure indicator, the auxiliary servo

buses, respectively. The dc system operation is automatic;

hydraulic pressure indicator, main gearbox oil pressure

control switches and rectifier caution lights are provided.

indicator, and the No. 2 engine torquemeter. The isolation

Normally, primary and monitor bus loads are assumed by

transformer is protected by a circuit breaker on the pilot

both rectifiers. Primary bus loads are those loads essential

circuit breaker panel marked XMFR NO.2 Ø C.

for flight under night instrument conditions and for

operation of equipment necessary for mission

2.6.1.7 Utility Ac Power Receptacle. There are one

requirements. The monitor bus loads are those not

or two

115-/200-vac utility receptacles on the sensor

essential for this type of flight. If one rectifier fails, the

operator console. There is one ac utility receptacle at the

associated reverse current cutout relay will remove the

No. 2 hoist operator station.

failed rectifier from the primary dc bus. The remaining

rectifier will assume the primary dc bus loads and the

monitored dc bus load will be dropped. The battery can

2-37

ORIGINAL

NAVAIR 01-230HLH-1

provide power to the primary dc bus when no other source

2.6.2.3 Battery. A

24-volt,

20-ampere hour, nickel

is available. The external dc power receptacle and

cadmium battery in the nose section forward of the pilot

associated circuitry permit use of an external power unit

compartment is reached from outside the helicopter.

for ground power application.

Battery power is used for limited ground operations when

no external power is available and as an emergency

2.6.2.1 Rectifers. Two

200-ampere,

28-vdc rectifiers

source of power to the primary dc bus. The transformer-

are incorporated in the system. The rectifiers require an ac

rectifiers supply charging current for the battery.

input from the generators or from an external ac power

source. The rectifiers are designated as Nos. 1 and 2, and

2.6.2.4 Battery Switch. The battery switch is on the

the associated components are designated in a similar

overhead switch panel (Figure

2-9) in the cockpit. The

manner. Both rectifiers normally supply power to the

switch is marked BAT and has marked positions OFF and

primary dc bus. The primary dc bus supplies power to the

ON. The ON position energizes the battery relay that

monitor dc bus. The No. 1 rectifier receives three-phase

connects the battery to the primary dc bus.

power from the No. 1 primary ac bus, and the No. 2

rectifier receives three-phase power from the No.

2.6.2.5 Utility Dc Power Receptacles. One utility dc

primary ac bus. The ac input is stepped down, rectified,

power receptacle is on the pilot overhead control panel

and filtered within each rectifier; the dc output is fed to

and another is on the right side of the cabin at the No. 2

the associated reverse current cutout relay. During normal

hoist operator station. These receptacles are powered by

operation, dc power from the reverse current cutout relay

the monitor dc bus and protected by two circuit breakers

is fed to the primary dc bus. The reverse current cutout

on the pilot circuit breaker panel under the general

relay prevents current flow from the primary dc bus to a

heading UT RECP and marked CKPT and CAB,

failed rectifier. The monitor dc bus will be dropped from

respectively.

the line if the ac monitor bus relay is open. The dc

monitor bus relay must be closed for the monitor bus to

2.6.2.6 External Dc Power Receptacle. The 28-volt

receive power. Power to close this relay comes from the

external dc power receptacle is on the right side of the

primary dc bus and is routed through the No. 2 and the

helicopter below the pilot window. External power can be

No. 1 reverse current cutout relays. If either the rectifier,

connected and used for all ground operation until the

reverse current cutout relay, or ac generator is inoperative,

generators are in operation. As soon as external power is

the monitor dc bus will be dropped from the line and the

connected, the external power relay is energized, external

appropriate caution light/lights will go on. The No.

power is supplied to the primary dc bus, the EXT PWR

rectifier is protected by a circuit breaker on the copilot

ON advisory light goes on, and the monitor bus relay is

circuit breaker panel marked RECTIFIER NO. 1. The No.

energized to permit the primary dc bus to furnish power to

2 rectifier is protected by a circuit on the pilot circuit

the monitor dc bus. The dc external power unit to be used

breaker panel marked RECTIFIER NO. 2.

for starting should provide

28 vdc,

300 amperes

continuous, and 750 amperes current limited.

2.6.2.2 Rectifier Caution Lights. Two rectifier

caution lights are on the caution light panel marked #1

RECTIFIER and #2 RECTIFIER. Failure of a rectifier or

reverse current cutout relay will light the associated

caution light.

2-38

ORIGINAL

NAVAIR 01-230HLH-1

2.7 LIGHTING EQUIPMENT

2.7.1.3 Console and Panel Lights. The lights on the

All lights operate on direct current and are protected

cockpit console, the overhead switch panel, the pilot right

by circuit breakers. Switches and rheostats for operating all

console, and the copilot interphone control panel are

lights, except, cabin dome light, flood-hover lights, and

controlled by a rheostat marked CONSOLE & PANEL

spotlight, are on the overhead switch panel.

LIGHTS with marked position OFF and BRT on the

overhead switch panel. The cockpit console and panel

2.7.1 Interior Lights

lights operate from the primary dc bus and are protected by

a circuit breaker marked CKPT under the general heading

2.7.1.1 Pilot and Copilot Flight Instrument Lights.

SWITCH PANEL on the center circuit breaker panel. On

The pilot and copilot flight instrument panel lights (Figure

UH-3H helicopters, the console and panel lights also

2-24) are individually controlled by rheostats marked

require power from the No. 2 primary ac bus. They are

PILOT FLIGHT INST LIGHTS and COPILOT FLIGHT

protected by a circuit breaker marked CKPT under the

INST LIGHTS, both with marked positions OFF and BRT

general heading PNL LTS CONV INPUT on the pilot

on the overhead switch panel (Figure 2-9). The intensity

circuit breaker panel.

of the flight instrument light may be varied by turning each

rheostat. When the pilot flight instrument light rheostat is

2.7.1.4 Instrument Emergency Light. This light is

moved out of the OFF position, it automatically dims the

controlled by the rheostat marked INSTRUMENT EMER

master reset light, caution panel, advisory panel, landing

LTS with marked positions OFF and BRT on the overhead

gear warning light, and fuel boost pump failure warning

switch panel. The intensity of the instrument emergency

lights. The pilot and copilot flight instrument lights operate

light that is a red light in the pilot compartment dome light

from the primary dc bus and are protected by circuit

(Figure 2-25) may be varied by turning the rheostat on the

breakers marked PLT and CO-PLT under FLT INST on

overhead switch panel. The instrument emergency light

the center circuit breaker panel and overhead control panel

operates on dc power, and it is protected by a circuit

(Figures 2-23).

breaker marked INST EMER on the center circuit breaker

panel.

2.7.1.2 Nonflight Instrument Lights. These lights on

the instrument panel are controlled by a rheostat marked

2.7.1.5 Pilot Compartment Dome Light. This light

NON-FLIGHT INST LIGHTS with marked positions OFF

(Figure

2-25) is controlled by a guarded switch marked

and BRT on the overhead switch panel. The intensity of

DOME LIGHTS - CKPT with marked p o s it i o n s RED,

the engine instrument lights, the hydraulic pressure gauge

OFF, and WHITE on the pilot compartment dome light

lights, the fuel management panel lights, the fuel quantity

panel. The dome light has a red and white lamp. The red

lights, and the engine and transmission oil pressure and

light may also be turned on by adjusting the instrument

temperature lights may be varied by turning the rheostat.

emergency light rheostat. The white light may be turned on

The nonflight instrument lights operate from the primary

only if the guard is lifted to permit moving the switch to

dc bus and are protected by a circuit breaker marked NON-

the white position. The white light in the pilot

FLT INST on the center circuit breaker panel and overhead

compartment dome light operates on dc power and is

control panel (Figure 2-23).

protected by a circuit breaker.

Figure 2-25. Pilot Compartment Dome Light Panel

2-45

ORIGINAL

NAVAIR 01-230HLH-1

2.7.1.6 Pilot Compartment Spotlights. Two

2.7.1.9

(ET) Ceiling Lights. The ceiling lights are

portable spotlights (Figure 2-24) with coiled cord are se-

mounted overhead in the cabin. The ceiling lights are

cured one on each side of the overhead switch panel. Two

controlled from either the forward bulkhead, STA 160, or

additional spotlights are installed, one on each side of the

aft bulkhead, STA 425, when the overhead control panel

compartment mounted on the window frame alongside

CABIN LTS MASTER switch is ON. Each station is

each pilot. The lights may be adjusted on their mountings

equipped with a pair of toggle switches. One switch for

to direct the light beams where required, or they may be

ON/OFF operation and the other for BRIGHT/DIM

removed and used as portable spotlights. The spotlights

operation. The lights receive power from the primary ac

are controlled either by rheostat with marked positions

bus and are protected by circuit breakers marked

OFF and BRT, or the pushbutton on the end of the

CEILING LTS fA and CEILING LTS fB on the

spotlight casing. The lens casing of the light may be

copilot’s circuit breaker panel.

turned to focus the beam and to position a red filter

converting the white light to a red light. The pilot

Note

compartment spotlights operate on dc power and are

protected by a circuit breaker.

On the UH-3H Executive Transport helicopters,

the BRIGHT/DIM switch on the CABIN

2.7.1.7 Cabin Dome Lights. These lights (Figure 2-

LIGHTS panel should remain in the DIM

24) are controlled by a guarded switch marked DOME

position while operating on external power.

LIGHTS - CABIN with marked positions RED, OFF, and

WHITE on the pilot compartment dome light panel

2.7.1.10

(ET) Reading. Reading lights are mounted

(Figure 2-25). The cabin dome lights have a red and white

overhead in the cabin. The CABIN LTS MASTER

lamp. The white light may be turned on only if the guard

switch must be ON to activate the individual pushbutton

is lifted. The intensity of the cabin dome lights may be

controls. The lights are powered by the primary dc bus

and are protected by a circuit breaker marked READING

varied by turning the rheostat marked CABIN DOME

on the overhead circuit breaker panel.

LIGHTS (Figure 2-26) at the sensor operator station. On

UH-3H helicopters, the panel contains two knobs marked

2.7.1.11

(ET) Instruction. The instruction lights

FWD and AFT, respectively, each with the marked

provide NO SMOKING and FASTEN SEAT BELT

positions DIM and BRT. The FWD knob controls the

information to the cockpit and cabin areas. In the cockpit

forward pair of dome lights, and AFT knob controls the

the information is displayed on the advisory panel. In the

aft pair of dome lights.

cabin area it is displayed on the forward bulkhead, STA

160, and aft bulkhead, STA 425. The information lights

are controlled by the PASSENGER INSTRUCTION

LIGHT switch located on the overhead control panel.

Circuit protection is provided by a circuit breaker marked

INST LTS on the overhead circuit breaker panel.

2.7.1.12

(ET) Emergency Exit. There are three

emergency exit lights. One located over each cabin

emergency escape hatch. The lights are operated by the

EMERGENCY LIGHTS switch. The lights receive

power from the primary dc bus and are protected by

circuit breakers marked EMER LTS 1 and EMER LTS 2.

Figure 2-26. Cabin Dome Light Panel

2.7.1.8 Cabin Panel Lights. These lights are con-

trolled by a panel at the hoist operator station, marked

CABIN PANEL LIGHTS with a knob marked DIM and

BRT to control light intensity. The cabin panel lights

operate from the primary dc bus and are protected by a

circuit breaker marked CAB under the general heading

SWITCH PANEL on the center circuit breaker panel.

2-46

ORIGINAL

NAVAIR 01-230HLH-1

2.7.2 Exterior Lights

2.7.2.1 Flood-Hover Lights

The flood-hover lights should not be on for

more than 15 minutes at a time to prevent

overheating. Allow a

10-minute cooling

period.

The floodlights

(Figure

2-24 and

2-24.1(ET)) are

on the electronics compartment door. A hover light is on

the lower leading edge of each stub wing. The flood-

hover lights are controlled by a switch (Figure 2-27) with

Figure 2-27. Flood, Hover, and Spotlight Control

marked positions HOVER LT, OFF, and FLOOD LT on

Switches

the pilot collective pitch lever grip. When the switch is

placed to FLOOD LT, the floodlights on the electronics

2.7.2.4 Anticollision Lights. These lights (Figure

compartment door light an area forward of the helicopter.

2-24), one on top of the tail pylon and the other on the

When the switch is placed to HOVER LT, the hover

bottom of the ARA-25A antenna, are controlled by two

lights on each wing section and the floodlights light an

switches. The aft two-position switch marked OFF and

area below and forward of the helicopter. Placing the

NORM on the overhead switch panel controls the forward

switch to the OFF centered position turns off the flood-

anticollision light. The forward three-position switch

hover lights. The flood-hover lights operate from the

marked BEACON, OFF, and ANTI-COLL on the

monitored dc bus and are protected by circuit breakers.

overhead switch panel controls the aft anticollision light

and beacon. When the forward switch is placed to the

2.7.2.2 Exterior Lights MASTER SWITCH. This

BEACON position, an aft white strobe light is activated.

switch marked MASTER SWITCH with marked positions

When the forward switch is placed to the ANTI-COLL

ON and OFF is on the overhead switch panel (Figure 2-

position an aft red strobe is activated. When the aft switch

9). The MASTER SWITCH must be placed ON before

is placed to the NORM position, the forward red

any of the position lights and rotating anticollision lights

anticollision light illuminates and rotates. The rotating

will operate. The tail beacon light may be operated with

anticollision light switch is inoperative until the exterior

the exterior lights MASTER SWITCH OFF.

light MASTER SWITCH is turned on. The forward

anticollision light will not operate unless the aft

2.7.2.3 Position Lights. The side position lights

anticollision light is turned on. The anticollision lights

(Figure 2-24) on the sponsons are controlled by a switch

operate from the primary dc bus. The control circuit is

marked SIDE POS with marked positions DIM, OFF, and

protected by a circuit breaker on the center circuit breaker

BRT on the overhead switch panel. The tail position light

panel. The anticollision lights are protected by circuit

on the tip of the tail pylon is controlled by a switch

breakers marked FWD and AFT under the general

marked TAIL POS with marked positions DIM, OFF, and

heading BEACON on the center circuit breaker panel. For

BRT on the overhead switch panel. The keyer operates

day operations, the white beacon is recommended; for

only when the position light switches are at either OFF or

night, ground, and shipboard operations, the red

DIM. A switch marked STEADY and FLASH is on the

anticollision lights are recommended.

overhead switch panel to permit automatic flashing of the

position lights. The position light switches are inoperative

2.7.2.5 Controllable Spotlight. Two spotlight control

until the exterior light MASTER SWITCH is turned on.

switches on the pilot collective pitch lever grip (Figure 2-

The position lights operate on dc power and are protected

27) control a swivel-type controllable spot-light (Figure

by a circuit breaker.

2-24) on the bottom right side of the fuselage aft of the

electronic compartment door opening. The left switch is

marked CONT SPOT LT with marked positions

MASTER, OFF, and RETRACT. The right switch is a

spring-loaded four-position thumb switch, center position

OFF, with marked positions EXTEND-RETRACT-L-R.

2-47

ORIGINAL

NAVAIR 01-230HLH-1

Placing the left switch to MASTER lights the controllable

continuity of the direct control linkage is maintained from

spotlight and furnishes power to the right switch to

the controls in the pilot compartment through the

control the spotlight. When the right switch is placed to

auxiliary and the primary servos to the rotary wing blades

EXTEND, the controllable spotlight is extended and may

except for a slight amount of end play at each servo unit

be stopped by releasing the switch to direct the light beam

to permit the pilot valves to move before the direct control

at any vertical angle between the stowed position to about

linkage. Normally, both servo systems are in operation at

30^o above the horizon. By placing the switch to

all times.

RETRACT, the light beam may be directed at a

progressively decreasing angle until the spotlight is in the

2.8.1 Primary Flight Control Servo System. The

fully stowed position. By placing the switch to L or R,

system consists of three hydraulic servo units that connect

the spotlight will turn to the right or left to any point in a

the flight control linkage to the stationary swashplate of

360^o arc. If the left switch is placed to RETRACT while

the rotor wing assembly. The servos provide the power

the controllable spotlight is extended, the spotlight will

necessary for operation of the rotary wing flight control

automatically go out and retract to the stowed position.

system only. The three servo units of the primary servo

The switch is then placed OFF. The controllable spotlight

systems are at the stationary swashplate. All three servo

operates from the primary dc bus and is protected by

units respond simultaneously and move in the same

circuit breakers.

direction in response to movements of the collective pitch

lever. Two of the servo units (lateral servo units) respond

2.7.2.6 Landing Gear Downlock Lights. These

simultaneously, but move in opposite directions in

lights

(Figure

2-24) on the main strut of each main

response to lateral movements of the cyclic stick. One of

landing gear provide additional landing gear position

the servo units (fore-and-aft servo unit) responds to fore-

information for an outside observer. Electrical power for

and-aft movements of the cyclic stick. Since all three

the lights is supplied by the primary dc bus through

movements can occur simultaneously through the action

contacts of the LDG GEAR DOWN LIGHT relay. The

of the mixing unit, the position of any primary servo unit

relay is energized whenever the exterior lights MASTER

is the result of the combined input of the cyclic stick and

SWITCH is placed in the ON position; however the

collective pitch lever. This results in a primary servo

circuit is not completed, thus preventing the lights from

system in which any one servo has an effect on both

going on until the landing gear is down and locked.

collective pitch and cyclic (lateral or fore and aft) pitch.

The primary servo hydraulic pump is driven by the

2.7.2.7 Droop Stop Light. The light (2-24) inside the

accessory section of the main gearbox. The primary

forward edge of the transmission fairing provides a means

hydraulic system reservoir (Figure

3-1), mounted aft of

of determining the position of the droop stops at night

the main gearbox, has a capacity of about 0.45 gallon of

during rotor shutdown. The droop stop light is controlled

hydraulic oil. The PRI SERVO PRESS caution light will

by the ROTOR HEAD LT switch with marked positions

go on when the primary servo pressure drops below 1,000

OFF and ON on the overhead switch panel. The light is dc

psi or is turned off.

powered and is protected by a circuit breaker marked

ROTOR HEAD on the center circuit breaker panel.

2.8.2 Auxiliary Flight Control Servo System. The

four servo units of the auxiliary servo systems are

2.8 FLIGHT CONTROL SERVO HYDRAULIC

between the mixing unit and the flight controls. Each

SYSTEMS

control input acts independently on the corresponding

servo. The rotary rudder pedals position the yaw servo.

The flight control servo hydraulic systems (FO-10)

The collective pitch lever positions the collective servo.

consists of a primary and an auxiliary flight control servo

The cyclic stick positions either the fore-and-aft servo, the

system. The servo systems are required for a power boost

lateral servo, or both. The auxiliary flight control servo

for the pilot to operate the controls. The servos also

system receives corrective signals from the ASE and

prevent feedback of vibratory loads to the control sticks

automatically introduces these into the flight control

and rudder pedals. Both servo systems operated from

system. It provides for controlled flight if the primary

independent hydraulic systems and both use similar servo

servo fails. The auxiliary servo hydraulic pump is driven

hydraulic units to vary the rotary wing and rudder blade

by the main gearbox accessory section. The auxiliary

pitch through the mechanical linkage of the regular flight

hydraulic reservoir

(Figure

3-1), aft of the primary

control system. Each servo unit consists of a bypass

hydraulic system reservoir, has a capacity of about 0.45

valve, sloppy link, power piston, pilot valve, and the ASE

gallon of hydraulic oil.

The AUX SERVO PRESS

valve in the auxiliary servo units only. The flight control

caution light will go on when the auxiliary servo pressure

system actuates the pilot valve that admits hydraulic oil

drops below 1,000 psi or is turned off.

into the servo unit. The servo output is connected to the

flight control linkage to provide the power boost. The

2-48

ORIGINAL

NAVAIR 01-230HLH-1

2.8.3 Flight Control Servo Switch. Both the primary

and auxiliary flight control servo systems are controlled

by the same three-position flight control servo switch on

the collective pitch grip lever (Figures

2-28 and 2-29).

The marked switch positions are PRI OFF and AUX OFF.

The flight control servo switch is protected by a dc circuit

breaker marked SERVO on the center circuit breaker

panel.

Both servo systems are normally in operation with

the switch in the unmarked center (ON) position. To turn

off the primary servo, the switch is placed to the forward

PRI OFF position; to turn off the auxiliary servos, the

switch is placed to the aft AUX OFF position. Stronger

cyclic, collective, and rotary rudder pedal forces and the

absence of pedal damping will be encountered with

auxiliary servos inoperative. The systems are

Figure 2-29. Copilot Collective Pitch Lever Grip

interconnected electrically in such a way that, regardless

of the switch position, it is impossible to turn either one

2.9 FLIGHT CONTROL SYSTEM

off unless there is 1,000 psi in the remaining system for

proper operation. The servo shutoff valves operate on dc

The flight control system is divided into three systems as

power.

follows: the rotary wing flight control system, the rotary

rudder flight control system, and the flight control

hydraulic power supply system with the three following

unique characteristics: (1) the collective to yaw coupling,

(2) the collective to the cyclic coupling, and

(3) the

negative force gradient installation. An ASE and coupler

system is installed, that, when engaged, provides

corrections of limited authority to the flight control

system to cause the helicopter to respond in a stable

manner to the maneuver referenced by the positions of the

flight controls. This equipment also provides automatic

cruising flight and constant altitude. The description and

operation of the ASE are included in the AUTOMATIC

STABILIZATION EQUIPMENT AND COUPLER

SYSTEM, paragraph

2.10. A beeper trim system is

installed to provide cyclic stick feel and to aid hands-off

control with the ASE and coupler system in operation.

2.9.1 Rotary Wing Flight Control System. This

system provides both vertical control and directional

control. Vertical control is accomplished by changing the

collective pitch of the rotary wing blades to increase or

decrease the angle of attack and consequently the lift

Figure 2-28. Pilot Collective Pitch Lever Grip

developed by the blades. Directional control is

accomplished by changing the pitch of each blade

2.8.4 Servo Hydraulic Pressure Indicators. The

individually as it rotates. The change in pitch causes the

primary and auxiliary servo pressure indicators operate on

blades to rise and fall as they rotate through 360^o, tilting

26 vac and are protected by circuit breakers. If either

the tip-path plane of the rotation of the rotary wing

servo system malfunctions, the malfunctioning system

blades, thereby obtaining a horizontal, as well as a

may be turned off and the helicopter flown on the other

vertical, component of thrust. The horizontal component

servo system. If the pressure in either the primary or the

of thrust moves the helicopter horizontally in whichever

auxiliary system drops below 1,000 psi, a pressure switch

direction the tip-path plane of rotation is tilted. Control

prevents the other system from being shut off regardless

motions from the collective pitch lever for vertical control

of the position of the servo switch.

and from the cyclic stick for directional control are

combined in a mixing unit in the ASE control

compartment aft of the pilot seat, and are transmitted to

2-49

ORIGINAL

NAVAIR 01-230HLH-1

the rotary wing assembly by mechanical linkage. Control

lowered. This provides attitude control during transitions

action is assisted by two hydraulically-operated flight

especially during ASE transitions while operating in

control servo systems. The rotary wing flight controls

coupler mode.

terminate at the stationary swashplate of the rotary wing

head. Control action is transmitted through the rotary

2.9.1.3 Collective Pitch Levers. Two collective pitch

swashplate and linkage on the rotary wing hub to the

levers (Figure 2-30) are in the pilot compartment, one to

blades.

the left of the pilot seat and the other to the left of the

copilot seat. Both levers operate simultaneously to change

2.9.1.1 Collective to Yaw Coupling. When the

the collective pitch of the rotary wing blades. A nut on the

auxiliary servo is pressurized, there is a proportional but

pilot collective pitch lever marked COLLECTIVE PITCH

irreversible transfer of collective pitch motion into the

LOCK with an arrow pointing left marked INCREASE

rotary rudder blade angle (collective pitch motion will act

FRICTION can be turned to apply friction to prevent the

to displace the rotary rudder but rotary rudder pedal motion

collective pitch lever from creeping while in flight.

will not affect rotary wing collective pitch blade angle).

This coupling provides automatic rotary rudder pitch

changes to compensate for collective pitch changes.

Rotary rudder blade angle changes result from both

WARNING

collective pitch lever and rotary rudder pedal inputs.

Any combination of collective pitch lever position

and rotary rudder pedal position, wherein the total would

The collective friction nut should not be

exceed the system limits, is nonattainable during flight.

completely disengaged from the friction

The collective pitch lever has overriding authority and

block, as this may result in the collective

therefore is always free to move within its full travel. If a

stick binding or jamming.

collective pitch lever position, added to the rotary rudder

pedal position, creates a rotary rudder blade angle equal to

the system limits, additional collective pitch lever motions

to exceed the limits is only possible at the sacrifice of

rotary rudder pedal position; the pedals will be forced to

move but the blade angle remains at the limit. With

auxiliary servo on, collective pitch lever low, and rotary

rudder pedal full left, raising the collective pitch lever to

high will be accompanied by pedal motion to the right.

With the collective pitch lever high and rotary rudder

pedal full right, reducing collective pitch lever to low will

be accompanied by pedal motion to the left. With the

auxiliary servo switch OFF, the operation is the same

except the irreversibility is not effective. Therefore, when

the combination of collective and pedal positions reaches

the system limit, additional collective motion is possible

by sacrificing pedal position. This trading of motion is not

Figure 2-30. Collective Pitch Lever

likely to occur but may be noted during ground check

with the auxiliary servo off. During rapid rotary rudder

pedal motions on the ground, noise can be heard aft of the

2.9.1.4 Cyclic Sticks. The cyclic stick in front of each

pilot seat provides directional control of the helicopter.

pilot seat when the pedals reach their right or left limits.

Moving the cyclic stick in any direction tilts the tip-path

The sound is created by the system stops and indicates

that the collective pitch and the pedals have reached the

plane of rotation of the rotary wing blades in that

direction and moves the helicopter in the same direction.

limits of the rotary rudder control. Additional rotary

The stick grip (Figure

2-31) contains pushbutton and

rudder pedal motion is possible by reciprocal motion of

the collective pitch lever.

thumb -operated switches for controlling various

equipment installed in the helicopter.

2.9.1.2 Collective to Cyclic Pitch Coupling. A bias

in the collective to cyclic pitch (fore and aft) coupling is

incorporated in the mixing unit to apply a nosedown

pitching correction automatically when the collective

pitch lever is raised and noseup when the collective is

2-50

ORIGINAL

NAVAIR 01-230HLH-1

Figure 2-31. Pilot and Copilot Cyclic Stick Grip

2.9.1.5 Force Gradient-Beeper Trim System.

trimmed reference point. The force gradient-beeper trim

The system permits a fine degree of adjustment of the

system will operate as long as there is both dc power and

cyclic stick and provides cyclic stick feel. When used with

auxiliary hydraulic pressure to the actuators. With the loss

the ASE engaged, the system permits hands-off flight by

of auxiliary hydraulic pressure only, the cyclic stick may

holding the stick in a selected trim position. The system

be repositioned by depressing either trim release button,

consists of two electrically controlled hydraulic actuators,

moving the cyclic to desired position, and releasing the

one mounted on the fore-and-aft channel of the auxiliary

trim release button. The cyclic will then be held in its new

servo and one on the lateral channel of the auxiliary servo.

position by hydraulic fluid trapped on both sides of the

Additional components of each trim actuator consist of the

actuator trim piston. In each case, the effect of the force

following: trim valve, force gradient spring, trim release

gradient spring will be felt when the cyclic is displaced

button, beeper trim button, and master beeper trim switch.

without pressing the trim release button. With system

The system is supplied with 28-vdc power and is protected

electrical failure, the pilot must manually override beeper

by a circuit breaker marked BEEPER TRIM on the center

trim pressure to reposition the cyclic.

circuit breaker panel. Hydraulic pressure for operation of

the system is supplied by the auxiliary hydraulic system at

2.9.1.5.1 BEEPER TRIM Switch. A switch marked

a reduced pressure of 60 psi. The cyclic stick may be

BEEPER TRIM-ON-OFF is on the overhead switch panel

manually displaced without disengaging the trim system

(Figure 2-9). The switch is the master control for the force

by applying about 1 to 1-1/2 pounds of pressure to the

gradient-beeper trim system and must be on before the

cyclic in the desired direction. A resistance force, caused

system will operate. When the switch is placed ON,

by force gradient spring compression, will increase about

hydraulic pressure holds the cyclic stick in position. With

1/2 pound for each additional inch of cyclic stick displace-

the beeper trim switch off, there is no force holding the

ment. When the pressure on the cyclic stick is released, the

cyclic stick. The pilot must manually control the cyclic.

spring compression will return the stick to its original

2-51

ORIGINAL

NAVAIR 01-230HLH-1

2.9.1.5.2 Beeper Trim. These switches on the pilot and

2.9.2.1 Rudder Pedals. The rotary rudder pedals (see

copilot cyclic stick grips

(Figure

2-31) have marked

Figure 2-3), one set in front of the pilot and the other in

positions FWD-AFT-L-R. This four-way thumb-operated

front of the copilot, change the pitch and thrust of the

switch is spring loaded to the center (off) position and

rotary rudder and consequently the heading of the heli-

when placed to any of the four positions, hydraulic

copter. Pressing the left pedal increases the rotary rudder

pressure will drive the cyclic stick in the selected direction.

blade pitch that increases thrust and turns the helicopter to

When the desired cyclic stick position is obtained, the

the left. Pressing the right pedal decreases rotary rudder

switch is released. The action of the force gradient system

blade pitch that decreases thrust and allows the helicopter

will then function about this location of the cyclic stick.

to turn to the right. Rotary rudder pedal adjustment knobs

are used to adjust the pedals for leg length. Electrical

2.9.1.5.3 Trim Release Button. The spring-loaded

switches mounted on the UH-3 pedals cancel directional

pushbutton switches are on the pilot and copilot cyclic

signals of the ASE when feet are placed on the pedals. Toe

stick grips (Figure 2-31) and are marked TRIM REL. The

brake pedals for the main landing gear wheelbrakes are

trim release buttons may be used to manually reposition

mounted on the pilot rotary rudder pedals.

the cyclic stick to a new reference point. Pressing either

switch applies dc power to the solenoids of both trim

Rotary rudder control pedal movements are

valves. The valves are then open and allow hydraulic fluid

transmitted to the rotary rudder by a series of mechanical

to pass through the return port. Both trim pistons may then

linkages and cables. Control inputs are transmitted from

move unrestricted within their cylinders. Releasing the

the pilot's and copilot's rudder pedals to the auxiliary servo

trim button deenergizes the solenoids, and the trim

by linkages and bellcranks and from the auxiliary servo by

actuators will then maintain the new cyclic stick reference.

two steel cables. These cables are routed through a series

of pulleys in the overhead of the cabin. Forward of the

2.9.2 Rotary Rudder Flight Control System.

tailfold hinge, the tail rotor control cables are attached to a

The functions of this system are to compensate for rotary

bellcrank that is attached to a series of control rods that

wing torque and to permit changing the heading of the

transmit control inputs to the tail rotor pitch beam (star).

helicopter. The torque developed by the rotary wing

blades turning counterclockwise tends to turn the fuselage

A rotary rudder NFG spring is installed parallel to the

in a clockwise direction. Gross weight, altitude, rate of

center pylon control rod at the top of the tail pylon just

climb, airspeed and the corresponding power settings, and

below the tail gearbox. The NFG spring is designed to

collective pitch will vary the amount of rotary wing torque.

relieve the pilot of rotary rudder forces created by

To compensate for torque variations, the pitch and

aerodynamic and inertia loads when the auxiliary hydraulic

resulting thrust of the rotary rudder blades can be increased

servo is inoperative, imparting a force into the flight

or decreased. Turns are accomplished by increasing rotary

control system against the rotary rudder aerodynamic

rudder thrust that overcompensates for rotary wing torque

forces.

and changes the heading of the fuselage to the left, or by

decreasing the rotary rudder thrust that undercompensates

Because of the rocker assembly bellcrank connecting

for the rotary wing torque and changes the heading of the

the NFG spring to the center pylon control rod, the spring

fuselage to the right. Rotary rudder control pedal

acts as an anticentering device that attempts to drive the

movements are transmitted to the rotary rudder assembly

tail rotor pitch and pedals away from the center (neutral

by mechanical linkage and cables. Control action is

pedal) position to maximum pitch in either direction. With

assisted by the auxiliary servo system only.

the main rotor static, when the auxiliary servo is secured

and the rudder pedals are displaced from neutral, the

A hydraulic damping device, incorporated in the yaw

spring will drive the tail rotor to either the full left or full

channel of the auxiliary servo, prevents abrupt movements

right pedal position depending on the direction of the

of the rotary rudder pedals. These movements would cause

initial pedal displacement.

sudden changes in thrust developed by the rotary rudder

with resulting rapid yaw acceleration and possible damage

2.9.2.2 Rudder Pedal Adjustment Knobs. These

to the helicopter. The rotary rudder pedal damper is

knobs are on each side of the cockpit, just forward of the

inoperative when the auxiliary servo system is inoperative

ashtrays. The adjustment knobs are connected to

or shut off. Yaw compensation is accomplished by

mechanical linkage that provide for fore-and-aft

mechanical linkage in the mixing unit that automatically

adjustment of the rotary rudder pedals. The knobs are

changes rotary rudder blade angles for changes in col-

turned to the right, as indicated by the arrow marked FWD,

lective pitch without moving the pedals, unless both

for forward adjustments and to the left, as indicated by the

collective pitch and rotary rudder blade angle are at their

arrow marked AFT, for the aft adjustment.

maximum limits, in which case the pedal will be forced back

with collective pitch change.

2-52

ORIGINAL

NAVAIR 01-230HLH-1

2.10

AUTOMATIC STABILIZATION EQUIPMENT

is placed in a synchronizing mode (no heading correction

(ASE) COUPLER SYSTEM

signal is developed) until his feet are removed from the

pedal switches. During the synchronizing mode, the yaw

The ASE provides added attitude and directional

rate gyro develops a signal proportional to the manual

stabilization and barometric hold. On UH-3H aircraft, an

heading displacement rate of the helicopter. This signal

SIU installed between the AHRS pitch and roll output and

initiates an open-loop spring condition that produces a

the ASE enhances transitions to automatic approach in the

proportional feedback force at the pedals. As the pilot

RAD ALT and the VERT ACCEL modes of the ASE

presses either pedal, he feels the proportional feedback

coupler. The coupler system is used with the ASE for the

force opposing the pedal pressure applied. The feedback

following: automatic approaches and doppler hovers using

force remains until the pilot has established the new

signals from the Doppler, radar altimeter, and

reference heading. Heading stability correction occurs

accelerometers. These signals are processed by the IVSC.

anytime the helicopter is displaced from the desired

The auxiliary servo system provides the means for

reference heading. In the collective channel, the altitude of

introducing ASE signals to the flight controls. UH-3H

the helicopter is held constant by signals developed from

Executive Transport helicopters do not use the Doppler or

the barometric altitude controller that senses changes in

sonar cable altitude signals. AC power supplies excitation

barometric pressure from the engage point. Automatic

and reference voltage for the ASE. The system is protected

barometric altitude stability correction occurs anytime the

by the three circuit breakers: two ac circuit breakers and

helicopter is displaced up or down from the reference

one dc circuit breaker.

altitude. The ASE may be engaged 3 minutes after power

has been applied to the system.

2.10.1 Automatic Stabilization Equipment. The

ASE improves the handling characteristics of the

2.10.1.1 ASE CONTROL Panel (BASIC ASE).

helicopter and permits hands-off automatic flight. The

The ASE CONTROL panel (Figure

2-32) is on the center

ASE may be engaged at all times, has less control

console.

authority than the primary flight control system, and may

be easily overridden through normal use of the flight

controls. The pilot has direct control of the ASE at all

times and can engage or disengage the entire system or any

channel as desired by means of switches on the ASE

CONTROL panel, CHANNEL MONITOR panel, cyclic

sticks, and collective levers. The hover indicators (UH-3H

helicopters) or flight directors

(UH-3H Executive

Transport helicopters) visually provide all ASE signals

into the pilot and copilot. The hover indicators visually

indicate all ASE signals to the pilot and copilot. The ASE

has attitude and directional stabilization and barometric

altitude hold modes of operation.

Attitude and directional stabilization are controlled

through the pitch, roll, and yaw channels; and barometric

altitude hold is controlled through the collective channel.

The ASE is capable of maintaining the barometric altitude

of the helicopter during normal flight, or when hovering

out of ground effect by using barometric altitude reference.

In the pitch and roll channels, the fuselage attitude is held

constant by comparing the actual attitude signal received

from the vertical gyro with the reference attitude trim

s ignal provided by the cyclic stick position sensor.

Automatic pitch and roll attitude stability correction occurs

anytime the helicopter is displaced from the trimmed

attitude. Pitch and roll gyro information source is selected

on the CHANNEL MONITOR panel. In the yaw channel,

Figure 2-32. ASE CONTROL Panel

the helicopter is held constant by comparing actual

heading signals received from the compass system with

reference heading signals received from the YAW TRIM

knob and the yaw synchronizer. While the pilot establishes

a reference heading by use of the pedals, the yaw channel

2-53

ORIGINAL

NAVAIR 01-230HLH-1

2.10.1.2 ASE Button. Pressing the ASE button will

The pitch and roll channels are identical in their

engage the pitch, roll, and yaw channels of the ASE and

operation.

The altitude channel has two modes of

the button will light. Once the ASE is engaged, the only

operation: radar altitude, and vertical accelerometer. The

noticeable difference in flight is that the helicopter is

vertical accelerometer mode is used when collective

dynamically stable.

pumping is experienced and for alternate approaches. The

ASE must be engaged before operating the coupler. Raw

2.10.1.3 Barometric Altitude Button. Pressing the

Doppler signals are used in the Doppler mode. When

BAR ALT button engages the barometric altitude

operating in the DOPP mode, the pilot selects the fore-and-

controller and the button will light. Engaging the altitude

aft groundspeed and drift groundspeed at which he desires

controller provides altitude hold. Once the BAR ALT

to fly. This is done through the use of the SPEED and

button is engaged, the pilots can momentarily release the

DRIFT set knobs, on the ASE CONTROL panel. The

altitude hold with the BAR REL button on both collectives

Doppler signals that are proportional to selected speed and

(Figures

2-28 and

2-29). For proper operation, it is

drift of the helicopter are compared to the output of these

important that BAR ALT be engaged at a zero vertical

controls and will maintain the helicopter at the selected

velocity and at the selected forward speed.

speed. The helicopter groundspeed and attitude are

controlled by the position of the cyclic stick, and the ASE

2.10.1.4 Barometric Altitude Off Button. Pressing

couplers operate around the cyclic stick position.

the BAR OFF button on the ASE control panel disengages

Therefore, if the cyclic stick is moved, the ASE and

the collective channel and the barometric altitude

coupler operating reference will move with it. If the

controller (BAR ALT) if the coupler is not engaged.

coupler is engaged and a coupled error signal exists, the

beeper circuit develops an electrical signal that actuates the

2.10.1.5 YAW TRIM Knob. The triangle-shaped YAW

appropriate hydraulic beeper valve to automatically

TRIM knob is used for small trim changes in forward

reposition the cyclic stick. It is through this operation that

flight and small turns while hovering without disengaging

the ASE maintains its authority.

the yaw channel. A large heading change with the YAW

TRIM knob during forward flight will result in a flat turn.

The radar altitude, ALTITUDE set potentiometer

For large heading changes, the pilot operates the cyclic and

controller knob (POT), and the Doppler vertical position

rotary rudder pedals in the normal manner.

signal (Vz) develop signals for the RDR ALT mode of

operation. The radar altimeter, ALTITUDE set POT, and

2.10.1.6 CG TRIM Knob. The cloverleaf-shaped CG

the vertical accelerometer (V_A) develop signals for the V_A

TRIM knob is used to compensate for changes in the

mode of operation. When operating in any of the coupler

center of gravity. The pilot should adjust the cg trim

modes, the pilot selects the altitude at which he desires to

whenever the hover indicator (NON-ET) or flight indicator

fly with the ALTITUDE set knob on the ASE CONTROL

(ET) indicates a need for adjustment. With the hover

panel. In the RDR ALT mode, the signal from the radar

indicator operating in the A mode and the METER

altimeter that is directly proportional to the absolute

SELECTOR switch on the CHANNEL MONITOR panel

altitude over the surface is compared to the output of the

in the ASE position, there may be a tendency for the pitch

ALTITUDE set knob, and the result will maintain the

bar to remain above or below the centerline. Such a condition

helicopter at the selected altitude.

The METER

indicates the need for cg adjustments. Maladjustment of the

SELECTOR switch on the CHANNEL MONITOR panel

CG TRIM knob reduces the efficiency of the pitch

(Figure

2-33) must be at ASE to present a proper

stabilization and may make the helicopter unstable in

indication.

pitch.

2.10.2.1

(NON ET) ASE CONTROL Panel

2.10.1.7 Automatic Stabilization Button. The AUTO

(Coupler). The automatic stabilization equipment must

STAB RELEASE button on the pilot and copilot cyclics

be engaged before engaging the coupler.

will disengage all ASE modes when depressed.

2.10.2.2

(NON ET) Coupler Button. Pressing the

2.10.2 (NON ET) Coupler System. The coupler

CPLR button engages the coupler; and the button lights

enables the helicopter to seek and retain selected absolute

whenever the coupler is engaged. Engagement of the

altitudes,

fore-and-aft

groundspeeds,

and drift

CPLR automatically engages BAR ALT. The BAR ALT

groundspeeds. The coupler will maintain a constant

cannot be disengaged with the BAR REL buttons on the

absolute altitude over the water. To do this, the coupler

collective sticks or with the BAR OFF button on the ASE

must control the helicopter in absolute altitude and the

CONTROL panel.

longitudinal and lateral axes. The pitch, roll, and altitude

channels of the coupler provide the signals that will control

the helicopter in the three axes needed for coupler

operation.

2-54

ORIGINAL

NAVAIR 01-230HLH-1

2.10.2.7

(NON ET) ALTITUDE Knob. The cross-

shaped absolute ALTITUDE set knob permits the pilot to

preselect accurate absolute altitudes from 0 to 200 feet

with the ASE and altitude coupler engaged. The

ALTITUDE set knob is scaled from 0 to 200 feet.

2.10.2.8 (NON ET) Hover Trim Engage Button. The

HOVER TRIM ENG button transfers doppler groundspeed

and drift control to the crewman at the hoist station for

positioning the helicopter during hoisting or rescue

operation. The HOVER TRIM ENG button lights when

engaged and can only be engaged when the CYC CPLR

switch is at DOPP.

2.10.3 Hover Indicators.

2.10.3.1 Pilot Hover Indicators. The pilot hover

indicators (Figure 2-34), mounted on the instrument panel,

is a four-axis indicator and has two modes of operation

that are selected by a knob on the face of the indicator. A

single OFF flag is used in both modes of operation. In the

A mode, it will retract when the ASE is engaged. In D

mode, the flag appears when the system is off or in

memory C mode is disabled.

2.10.3.2 A Mode. The A mode monitors the ASE or

coupler, depending on the position of the METER

SELECTOR switch. With the METER SELECTOR

switch on the CHANNEL MONITOR panel in the ASE

Figure 2-33. Channel Monitor Panel

position, the hover indicator will operate as a null

indicator, indicating the input to the ASE servo valves. The

2.10.2.3 (NON ET) Cyclic Coupler Switch. The CYC

hover indicator horizontal bar is used to monitor the pitch

CPLR switch OFF position makes it possible to disable the

channel; the vertical bar is used to monitor the roll

cyclic coupler even though the altitude coupler is working.

channel; the vertical pointer is used to monitor the altitude

The DOPP position selects the Doppler radar input for

channel; and the horizontal pointer is used to monitor the

automatic cruise, transition, and hover.

yaw channel. The hover indicator scale factor in this mode

is a

2-milliampere per division of servo valve differential

2.10.2.4 (NON ET) Altitude Coupler Switch.

current with 8 milliamperes considered as a hardover. The

The RAD ALT and VA positions of the altitude

A mode with the METER SELECTOR switch at CPLR, is

coupler switch engage the appropriate altitude channel

used to monitor coupler output.

sensors for absolute altitude reference.

The VERT

ACCEL position engages the vertical accelerometer mode

2.10.3.3 D Mode. The D mode connects the hover

to eliminate collective pumping over high sea states. The

indicator to the Doppler. In this mode, the horizontal bar

CYC CPLR need not be engaged in order to engage the

indicates fore or aft velocity, the vertical bar indicates left

RAD ALT or VA modes of the altitude coupler.

or right drift, the vertical pointer indicates vertical

velocity, and the horizontal pointer is inoperative. The

2.10.2.5

(NON ET) DRIFT Knob. The bar-shaped

scale factor for the D mode is 10 knots per division on

DRIFT groundspeed set knob permits the pilot to preselect

horizontal and vertical bars for groundspeed. To maintain

the lateral drift of the helicopter to be maintained by the

a Doppler hover, the pilot can assume the cyclic stick is in

cyclic coupler.

the center of the indicator. If the horizontal bar moves

upward, the pilot should move the cyclic toward the bar or

2.10.2.6 (NON ET) SPEED Knob. The indented circle-

in a forward direction. The pilot flies the helicopter to the

shaped, fore-and-aft ground SPEED set knob permits the

bars. The vertical pointer can be used as a vertical velocity

indicator with each division equal to 250 fpm.

pilot to preselect fore or aft groundspeed to be maintained

by the cyclic coupler.

2-55

ORIGINAL

NAVAIR 01-230HLH-1

2.10.3.4 (ET) Flight Director

The flight director (Figure

selected course is a bearing to or from a surface navigation

2-34.1), on the instrument panel, allows the pilot and/or copilot

beacon. When the knob is at T, power for warning flag

to visually monitor composite information from the ASE, VOR

operation is supplied from the TACAN set. Exact bearing

#1, and VOR #2 receivers, glideslope receiver, and TACAN

information is available on the TACAN course indicator that is

set. It has four modes of operation that are determined by the

connected in parallel with the flight director. Since there are

position of the mode selector knob. The pilot and copilot can

two completely independent VOR receivers, separate operating

monitor the same mode or any combination of modes by

modes are available to the flight director. When the VOR #1

placing either knob in the A, T, V1, ILS, or V2 position. The

receiver is turned on and the knob is at V1, ILS, the indicator

mode indicator window in the upper right quadrant of the face

monitors lateral deviation from a VOR course that was selected

indicates ASE, TAC, VOR 1, ILS, or VOR 2 depending upon

on the VOR #1 course indicator, and longitudinal deviation

which mode has been selected. With ASE engaged and the

from a corresponding glideslope signal. The top center

knob at A

(ASE visible in mode indicator window), the

warning flag on the flight director disappears when the radio is

indicator monitors output signals from the pitch, roll,

turned on and the bearing signal received is reliable. The right

collective, and yaw ASE channels. Each output signal

middle warning flag disappears when the glideslope signal

represents the amount of differential current flow to the servo

received is reliable. The flight director does not indicate if the

valve in a particular channel. The warning flags disappear

lateral position deviation from a selected course is a bearing to

when ASE is engaged. During ASE mode of operation, all

or from an omnidirectional range source. Power for the

meter movements are used. The horizontal bar monitors signal

warning flag during VOR #1 operations is supplied from the

output from the pitch channel and indicates forward or aft

VOR #1 receiver. Exact bearing information is available on

displacement from selected attitude. The vertical bar monitors

the VOR #1 course indicator that is connected in parallel with

signal output from the roll channel and indicates left or right

the flight director. With the VOR #2 receiver turned on and the

displacement from selected attitude.

The vertical arrow

knob at V2, the entire operation is identical to VOR

monitors signal output from the collective channel and

operation except that glideslope information will not be

indicates up or down displacement from selected attitude. The

displayed

horizontal arrow monitors signal output from the yaw channel

and indicates left or right displacement from selected direction.

2.10.4 HOVER TRIM Control Panel. The panel is

All ASE signal inputs to the indicator are routed through the

mounted just forward of the cabin door. The hover trim stick is

channel monitor control panel. Power for warning flag

mounted on the panel in the center of a diamond-shaped outline

operation is supplied from the AUTO STABE circuit breaker

marked R (right), L (left), F (forward), and A (aft) directional

through the ASE ENG button. The warning flag in the top

control. To use the hover trim stick, the coupler must be

center of the face indicates when the indicator is operating or is

functioning in the DOPP mode and the HOVER TRIM ENG

turned off for a particular mode of operation. The flight

depressed. The red engage light on the tip of the stick will go

director has four-meter movements consisting of two bars and

on when the hover trim controls are ready for use. Two knobs

two arrows. The horizontal bar deflects up and down along a

on the panel are marked ROLL BIAS and PITCH BIAS. The

straight line of parallel reference marks on the vertical axis.

ROLL BIAS knob has direction-indicating arrows marked

The vertical bar deflects left and right along a straight line of

RIGHT and LEFT and the PITCH knob has arrows marked

parallel reference marks on the horizontal axis. The horizontal

FWD and AFT. The use of these knobs permits trimming for an

and vertical reference marks intersect and form a cross within a

accurate hover when the hover stick is centered by permitting

circle. This intersection point is the center, or null, of the

the aircrewman to introduce signals into the amplifier to correct

horizontal and vertical bars. The vertical arrow deflects up and

any slight drifting while hovering in the Doppler mode.

down along a straight line of parallel reference marks that are

perpendicular to the vertical reference edge. The horizontal

2.10.4.1 Hover Trim Stick. The hover trim stick (Figure 2-

arrow deflects left and right along a straight line of parallel

35) has been incorporated to give the crewman longitudinal and

reference marks that are perpendicular to the horizontal

lateral control of the helicopter. The 6-inch stick is mounted on

reference edge. The reference marks (divisions) are speeds

the HOVER TRIM control panel and is used in the same

equal to 0, 25, 50, 75, and 100 percent of full-scale deflection

manner as the cyclic stick. The HOVER TRIM control panel

in either direction. When the TACAN set is turned on and the

allows the operator to control the helicopter drift from 0 to 11

mode selector knob is at T (TAC), the Indicator monitors

±4-1/2 knots. When the hover stick is centered, the ROLL

lateral deviation from the selected course, relative to a surface

BIAS and PITCH BIAS knobs allow trimming for an accurate

navigation beacon. The center warning flag disappears when

hover by allowing the hoist operator to induce signals into the

the TACAN set is turned on and the bearing signal received is

amplifier to correct any slight drifting of the helicopter while

reliable. During TAC mode of operation, only the vertical bar

hovering in the Doppler mode. The hover trim stick can only be

is used. The vertical bar monitors signals received from the

engaged by the pilot during the Doppler mode of operation.

indicator coupler and indicates lateral position deviation from

the selected course on the TACAN course indicator. The flight

director does not indicate if the lateral position deviation from a

2-57

ORIGINAL

NAVAIR 01-230HLH-1

starboard gyro for ASE pitch and roll reference; however,

PORT is the preferred position. Electrical power for the

CHANNEL MONITOR panel is supplied from the primary bus

and is protected by a circuit breaker.

2.10.6 Channel Monitor Test Switch. The CHAN MON

test switch on the pilot compartment dome light panel (Figure

2-25) has two marked positions, TEST and OFF. OFF disables

all hardover switches but does not affect any of the other

switches. TEST permits use of the hardover switches by

maintenance personnel for testing.

2.10.7 Vertical Gyros. The 1080Y starboard gyro always

provides pitch and roll information to the pilot vertical gyro

indicator and also to the ASE when the VERTICAL GYRO

switch on the CHANNEL MONITOR panel is at STBD. The

compass system always provides pitch and roll information to

the copilot attitude indicator and also to the ASE when the

VERTICAL GYRO selector switch is at PORT. Power failure

flags will be displayed on the pilot attitude indicators whenever

a power failure is sensed or an imbalance in the ac power

applied to the vertical gyros is sensed.

2.11 UTILITY HYDRAULIC SYSTEM. The utility hydraulic

system (FO-11) provides hydraulic pressure for all hydraulic

Figure 2-35. Hover Trim Stick

equipment not included in the flight control servo and ASE

systems. The utility hydraulic system reservoir (Figure 3-1), aft

2.10.5

CHANNEL MONITOR Panel. The

CHANNEL

of the main gearbox, supplies hydraulic fluid to the utility

MONITOR panel (Figure 2-33) on the pilot console has four

system. In addition, the rotor brake hydraulic cylinder is

ASE channel disengage switches mounted across the top of the

gravity-feed with hydraulic fluid from the utility reservoir. The

panel. Directly beneath the channel disengage switches are four

utility hydraulic reservoir has a capacity of

1.09 gallons of

associated guarded hardover switches marked PITCH, ROLL,

hydraulic oil. The utility hydraulic pump on the accessory drive

COLL, and YAW.

section of the main gearbox provides

3,000-psi hydraulic

pressure. Equipment operated by the utility hydraulic system

2.10.5.1 Channel Disengage Switches. These switches

includes the main landing gear, rescue hoist, and fold system.

with marked positions ON and OFF allow the pilot to

On UH-3H Executive Transport the rescue hoist and sonar

disengage individual channels as desired.

reeling machine have been removed.

2.10.5.2 Hardover Switches. These switches permit

2.11.1 Utility Hydraulic Pressure Indicator. The utility

maintenance personnel to introduce hardover signals to

hydraulic pressure indicator

(Figure

2-7) on the instrument

individual channels of the ASE for testing. The hardover

panel operates on 26-vac power and is protected by a circuit

PITCH switch has marked positions FWD and AFT; the

breaker. The gauge marked UTI indicates pressure in the utility

hardover ROLL switch has marked positions LEFT and

hydraulic system in psi.

RIGHT; the hardover COLL switch has marked positions UP

and DOWN; and the hardover YAW switch has marked

2.12 LANDING GEAR SYSTEM. The landing gear system

positions LEFT and RIGHT.

consists of sponsons, retractable main landing gear assemblies,

a hydraulic system, and a fixed tailwheel. The main landing

2.10.5.3

(ET) METER SELECTOR Switch. The METER

gear consists of dual wheels, equipped with hydraulic brakes,

SELECTOR switch permits either ASE or CPLR inputs to be

that are attached to the sponsons by retractable oleo shock

monitored on the hover indicators

(UH-3H helicopters) or

struts. The main landing gear is equipped with a one-shot

flight directors (UH-3H Executive Transport helicopters) when

pneumatic emergency blowdown feature that permits lowering

the mode selector knob is in the A mode

the main landing gear. The main landing gear hydraulic system

(FO-12) operates on

3,000-psi hydraulic pressure from the

2.10.5.4 VERTICAL GYRO Selector Switch. The

utility hydraulic system. The system is actuated by dc electrical

VERTICAL GYRO selector switch with marked positions

power and is protected by a circuit breaker (marked LAND

PORT and STBD allows the pilot to select either the port or

GEAR on the center circuit breaker panel). The tailwheel

2-58

ORIGINAL

NAVAIR 01-230HLH-1

beneath the tail fin is full swiveling and self-centering and may

The main landing gear warning light in the landing gear

be locked in the center position. The sponsons are fixed,

actuating lever knob goes on whenever the actuating handle is

hollow, outrigger-type floats attached to the fuselage that

moved and the landing gear is in transit to the up or down

enable the helicopter to maintain a level, upright position in the

position. The intensity of the landing gear warning light is

water. The landing gear control panel (Figure

2-36) marked

controlled by the pilot flight instrument lights rheostat on the

LDG GEAR CONT is on the copilot side of the cockpit

overhead switch panel. When the rheostat is OFF, the light

console. The landing gear actuating lever, the warning light test

will be on at full strength, but as the rheostat is moved from

button marked HDL LT TEST, and the downlock release

OFF, the landing gear warning light will go to the dim position.

marked DN LCK REL are on the panel.

When the landing gear is locked in either UP or DOWN, the

light will go off. The warning light operates on dc power and is

2.12.1 Landing Gear Actuating Lever. The landing gear

protected by a circuit breaker marked LAND GEAR on the center

actuating lever on the landing gear control panel marked LDG

circuit breaker panel.

GEAR CONT has a wheel-shaped knob and marked positions

UP and DN, with directional arrows. The lever is actuated to

raise or lower the main landing gear and to energize or

deenergize the armament release. When at UP, the armament

release is energized, and when at DN, the armament release is

deenergized. A red warning light in the wheel-shaped knob is

on when the landing gear is cycling. An electrically actuated

downlock solenoid locks on the actuating lever in the DN

position when the weight of the helicopter is on the landing

gear. After becoming airborne, the lock is automatically

released that permits the actuating handle to be moved to UP.

Should the downlock solenoid electrical circuit become

inoperative, a mechanical downlock release marked DN LCK

REL can be actuated to mechanically release the landing gear

actuating lever from the DN position. Placing the actuating

lever to UP opens an electrically operated solenoid valve,

allowing hydraulic fluid to pass through a two -way restrictor

valve to the main landing gear cylinder, forcing the piston in

the cylinder into the up position. The main landing gear, when

fully raised, is held in position by a mechanical uplock. A

hydraulically operated uplock cylinder or a mechanical release

must be a actuated to release the mechanical uplock. Placing

Figure 2-36. Landing Gear Control Panel

the actuating lever to DN opens an electrically operated

solenoid valve, allowing hydraulic fluid to release the

2.12.3.1 Landing Gear Warning Light Test Button. A

mechanical uplock and to force the piston in the main landing

landing gear actuating lever warning light test button marked

gear cylinder to the down position. A limit switch will be

HDL LT TEST is on the landing gear control panel on the pilot

actuated when the main landing gear is fully extended or

compartment console. Pressing the button tests the warning

retracted and the electrically operated solenoid valve will return

light in the landing gear actuating lever knob. The warning

to the trail position.

light test button operates on dc power and is protected by

circuit breakers.

2.12.2 Down-Limit Release

(Scissors) Switch. One

down-limit release switch is attached to the scissors of each

2.12.4 Landing Gear Position Indicators. The indicators

main landing gear. With the weight of the helicopter on the

are on the cockpit console. The indicators read UP only if the

wheels, the main landing gear strut is compressed and the

landing gear wheels are in the up and locked position, and

scissors switch is electrically open. As the main landing gear

show pictures of landing gear wheels only if the wheels are in

strut extends, the scissors switch closes and the 28-vdc circuit

the down and locked position. During landing gear extension or

is completed, allowing the down-lock solenoid to disengage,

retraction (when the landing gear is neither up and locked nor

disabling underfrequency protection for the aircraft electrical

down and locked) and whenever electric power is not available,

system and enabling the jettison control panel.

the indicators show black and white diagonal lines. The

indicators operate on direct current and are protected by a circuit

2.12.3 Landing Gear Warning Light. When the landing

breaker marked LAND GEAR on the center circuit breaker

gear is fully extended, the red indicator pin in the center of the

panel.

drag link should be in.

2-59

ORIGINAL

NAVAIR 01-230HLH-1

2.12.5 Landing Gear Emergency System. The system

2.12.5.2 Emergency Landing Gear Release Lever. The

is used to lower the main landing gear pneumatically if the

lever

(Figure

2-38) marked EMERGENCY LANDING GEAR

hydraulic system fails. The landing gear emergency system

RELEASE PULL FWD, ONLY IF AIR RELEASE FAILS is

consists of a 50-cubic inch, 3,000-psi maximum capacity air

behind the pilot seat on the broom closet. Pulling the

bottle in the auxiliary servo controls enclosure (broom closet)

emergency landing gear release lever manually actuates the

immediately aft of the pilot, an emergency landing gear

emergency bypass valve, allowing hydraulic fluid trapped in

extension handle on the copilot side of the cockpit console, and

the uplines to be vented to the hydraulic system. Gravity will

an emergency landing gear release lever on the side of the auxiliary

then lower the landing gear to the down position. The emer-

servo controls enclosure. A gauge, visible through a window on

gency landing gear release lever should be pulled only after the

the control enclosure door indicates the nitrogen pressure in the

emergency landing gear extension handle has been pulled and

air bottle. Nitrogen pressure must be 2,500 to 3,000 psi for the

the landing gear still does not lower. Do not attempt to reset the

system to function. An air bottle filler cap is behind a hinged

lever after it has been placed to the forward or emergency

panel marked EMERGENCY LANDING GEAR RELEASE

position.

AIR CHARGING CONNECTION on the side of the auxiliary

servo controls enclosure.

2.12.5.1 Emergency Landing Gear Extension Handle.

The handle (Figure 2-37) painted with orange-yellow and black

diagonal stripes is on the copilot side of the cockpit console.

The emergency landing gear extension handle is used to lower

the main landing gear in case of failure of the normal system.

The handle must be turned and pulled to withdraw the

emergency uplock release pins, release air into the landing gear

system, and pneumatically actuate the emergency bypass valve

to vent hydraulic fluid from the uplines, forcing the landing

gear to the down and locked position. The emergency bypass

valve displaced by the air charge and actuated by the

emergency landing gear extension handle must be manually

reset before the landing gear can be retracted or the air bottle

recharged. An instruction plate marked EMER L.G.

EXTENSION, TURN THEN PULL is beside the handle. When

the emergency landing gear release system has been actuated,

Figure 2-38. Emergency Landing Gear Release Lever

the warning light on the landing gear actuating lever will go on

whenever the landing gear is in transit, regardless of lever

2.12.6 Tailwheel Lock Handle. The handle, next to a decal

position. The landing gear position indicator will indicate down

marked TAILWHEEL LOCK - PULL TO LOCK, is on the

or unsafe only.

right side of the cockpit console. Pulling the handle out permits

a spring-loaded lockpin to engage at the swivel joint after the

tailwheel is centered. Pushing the handle in releases the lock

and allows the tailwheel to swivel. There are positive detents in

the LOCK and UNLOCK positions. To UNLOCK the tail-

wheel, depress the button in the center of the handle to release

the handle from the upper detent, then depress the handle fully

down. It is not necessary to continue to depress the center

button, as the handle will automatically remain in the down

position when it engages the lower detent. The tailwheel

maybe unlocked before taxi since the tailwheel will not

actually unlock until all side loads on the lockpin are relieved.

To lock the tailwheel, depress the button in the handle. This

will allow the handle to rapidly spring out to the locked

position. The tailwheel should be locked only for straight

takeoffs and landings. During maneuvers on the ground, the

tailwheel should be unlocked to reduce strain on the pylon and

Figure 2-37. Emergency Landing Gear Extension Handle

the possibility of shearing the tailwheel lockpin. If any side

loads are imposed on the tailwheel, the lockpin will not

disengage.

2-60

ORIGINAL

NAVAIR 01-230HLH-1

2.12.7 Wheelbrake System

allowing the pilot to see the marking CHECK BLADE FOLD. The

system is protected by a circuit breaker marked BLADE FOLD

on the center circuit breaker panel.

2.13.1 Rotary Wing Blade Folding. The rotary wing blades

may be folded by a full automatic operation that is controlled

by either the pilot or copilot. The operation may be halted at

any stage of the folding or spreading cycle and reversed if so

The parking brake can be set for one wheel

desired. Hydraulic power from the utility hydraulic system is

without the other which could be hazardous

used in conjunction with hydraulic and electrical sequencing.

during night carrier operations. Make sure that

Power is supplied by a variable delivery pump on the accessory

both rudder pedals are pressed firmly when

drive of the main gearbox. The folding is done by actuating the

setting the parking brake.

proper switches on the blade fold control panel (Figure

2-39)

on the overhead switch panel and observing the blade

Note

positioning and folding in conjunction with appropriate lights

on the panel. The No. 1 engine should be running at 104-

The advisory light only indicates that the parking

percent N_f for blade folding. The No. 1 blade positions directly

brake handle is up. It does not infer the

aft and does not fold. After the rotary wing blades are folded

wheelbrake system is pressurized.

and stowed, the tail pylon may be manually folded without any

interference. The blade positioner control valve is a four-way

The main landing gear wheels have hydraulic brakes

trail position solenoid-operated valve mounted on the upper

operated by toe pedals on the pilot rotary rudder pedals. The

right side of the main gearbox input housing. It controls the

fluid for these brakes is self-contained within the master

operation of the positioning units. When the folding switch is

cylinders. A parking brake handle operates a hydraulic valve to

placed to FOLD, a solenoid energizes the control valve that

lock the wheel-brakes. A light on the advisory panel goes on

directs hydraulic fluid to engage the rotor brake disc teeth and

whenever power is applied to the helicopter, allowing the pilot

position the No. 1 blade aft. After positioning the blades, the

to see the marking PARKING BRAKE ON.

control valve is energized by the other solenoid to direct

hydraulic fluid to disengage the rotor blade positioners and,

2.12.8 Brake Pedals. The main landing gear wheels are

through a pressure reducer, to engage the automatic rotor

individually braked by pressing the corresponding toe brake

brake. When the folding switch is placed to SPREAD, the

pedals (Figure 2-3) mounted above the pilot rotary rudder pedals.

control valve is energized to direct hydraulic fluid to ensure

The brakes operate on hydraulic pressure developed by

disengagement of the positioner from the rotor brake disc and,

pressing the brake pedals.

through a pressure reducer, to engage the rotor brake.

2.12.9 Parking Brake Handle. The handle (Figure

2-3)

Note

marked PARKING BRAKE is on the right side of the cockpit

console. A decal next to the parking brake handle is marked

A manual blade fold procedure is described in

ON-DEPRESS TOE BRAKE THEN PULL, OFF-DEPRESS

Chapter 7.

TOE BRAKE. The parking brake is applied by first pressing

the toe brake pedals and then pulling the parking brake handle

out. Pressing the left brake pedal will release the parking brake,

causing the parking brake handle to return to OFF.

2.13 AUTOMATIC BLADE FOLD SYSTEM

The automatic blade fold system positions the rotary wing

blades into the proper radial position for blade folding, extends

the dampers and turns the blades a full 3^o forward on the drag

hinges, locks the rotary wing controls when the blades are

folded, provides a means of applying the rotor brake auto-

matically, and automatically disengages the blade lockpins.

Controls for the system are grouped on a panel marked MAIN

BLADE FOLD on the cockpit overhead switch panel. The

blade fold system is a part of the utility hydraulic system. A

Figure 2-39. Blade Fold Control Panel

light on the advisory panel goes on when dc power is applied

and the blade fold interlock system is not in flight condition,

2-61

ORIGINAL

NAVAIR 01-230HLH-1

2.13.2 Blade Fold Control Panel. The panel (Figure 2-

locking pin is properly positioned, safety valve is closed, and

39) marked MAIN BLADE FOLD on the overhead switch

master switch is off. When the safety valve switch is open and

panel contains all control switches for the blade fold system

the master switch is on, the flight position indicator light will

and the warning and indicator lights that show the sequence of

go off, SAFETY VALVE and FOLD POWER lights will

operation during the folding and spreading of the rotary wing

illuminate, primary pressure will drop to zero, primary servo

blades. Three switches are marked MASTER, SAFETY

caution light will go on, automatic rotor brake will come on,

VALVE, and BLADES with marked positions FOLD,

and rotor brake caution light will remain on.

SPREAD, and OFF. The indicator lights are marked BLADES

SPREAD, FLIGHT POS, NO. 1 BLADE POS, CONT LOCK

2.13.2.6 No. 1 Blade Position Indicator Light.

PINS ADV BLADES FOLDED, and PYLON UNLOCKED.

An amber light marked NO. 1 BLADE POS on the blade fold

Two warning lights are marked SAFETY VALVE OPEN and

control panel will go on during the folding cycle after the

FOLD PWR ON. The lights may be dimmed by turning the caps.

rotary wing head positions with the No. 1 blade directly aft.

2.13.2.1 SAFETY VALVE Switch. A guarded switch

2.13.2.7 Control Lockpins Advanced Indicator Light.

marked SAFETY VALVE with positions OPEN and CLOSED

An amber light marked CONT LOCK PINS ADV on the blade

is on the blade fold control panel. A red warning light to the

fold control panel will go on during the folding cycle as soon as

left of the safety valve switch indicates the OPEN position. The

one flight control lockpin moves forward. It will remain on

switch operates a motor-driven, three-way, two-position,

until all the flight control lockpins are fully disengaged at the

selector valve on the bulkhead on the right side of the cabin

end of the spreading cycle.

next to the gearbox and prevents inadvertent application of

hydraulic pressure to the blade fold system and automatic rotor

2.13.2.8 Blades Folded Indicator Light. An amber light

brake during flight. When the safety valve switch is turned to

marked BLADES FOLDED on the blade fold control panel

OPEN, the safety valve red warning light will go on and the

will go on when all blades are folded. During the spreading

FLIGHT POS green indicator light will go off.

cycle, the BLADES FOLDED indicator light will go off when

any blade moves forward. The blade fold accumulator located

2.13.2.2 Blade Fold MASTER Switch. The switch marked

in the No. 1 main rotor blade sleeve spindle is a piston-type

MASTER with two positions ON and OFF on the blade fold

oleo-pneumatic system designed to maintain hydraulic pressure

control panel completes the circuit to furnish electrical power

in the blade fold system when the blades are folded and power

to actuate the hydraulic sequences for automatic blade folding.

is off the system. By maintaining pressure on the system, the

The master switch will not furnis h electrical power to the blade

blade damper-positioners are maintained in the extended

fold system to fold the blades if the pylon is folded, accessory

position, locking the blades in the folded position. Addition-

drive switch is in FLIGHT, No. 2 engine fuel firewall valve is

ally, the blade fold accumulator compensates for expansion and

open, or safety valve is closed. A red indicator light marked

contraction of trapped hydraulic fluid because of temperature

FOLD PWR ON indicates that the folding power is available.

changes and helps dampen out pressure surges during the

fold/spread cycle. A pressure gauge on the No. 1 blade sleeve

2.13.2.3 Blades FOLD Spread Switch. A three-position

spindle indicates accumulator pressure. Normal pressure

BLADES switch marked FOLD, SPREAD, and OFF on the

indication is 1,500 psi with the blades spread.

blade fold control panel is used to select the folding cycle that

consists of rotary wing head positioning cycle, blade folding

2.14

PYLON FOLDING

cycle, or the spread cycle.

The tail pylon may be folded forward, parallel to the fuselage,

so the helicopter can be parked in a small area. The pylon is

2.13.2.4 Blades Spread Indicator Light. An amber light

attached to the tail cone at four points just forward of the

marked BLADES SPREAD on the blade fold control panel is

intermediate gearbox. The two right-hand points form a hinge

actuated when the blade lockpins are engaged and the control

(Figure 2-2) about which the pylon folds forward against the

lockpins are completely disengaged. The blade spread indicator

right side of the tail cone. A locking strut holds the pylon in the

light will go off when the accessory drive switch is placed to

folded position (Figure 2-40). A pylon hinge lock locks the left

FLIGHT. The BLADES SPREAD light will go off as soon as

side of the pylon to the tail cone in the spread position.

any control lockpin starts to advance or any blade lockpin

Retractable lockpins are actuated in and out of hinge-like lugs

retracts.

by a ratchet wrench that is secured against the tail cone when

not in use. The rotary rudder pedals should be centered before

2.13.2.5 Flight Position Indicator Light. A green light on

the pylon is folded to assure minimum rotary rudder pitch

the blade fold control panel marked FLIGHT POS indicates

change as the pylon is swung forward. A red lockpin position

that the rotary wing blades are in the flight position. The flight

flag-type indicator, operated mechanically, extends from the

position indicator light will go on, primary hydraulic pressure

left side of the tail cone whenever the pylon lockpins are not

will rise to 1,300 to 1,600 psi, and the primary servo caution

fully seated.

light will go off when the blades are completely spread, pylon

2-62

ORIGINAL

NAVAIR 01-230HLH-1

2.15.5 Attitude and Heading Reference System

(A/A24G-39 AHRS). This system provides PORT GYRO

attitude information (pitch and roll) for display on the copilot

instrument panel, attitude indicator, and the ASE when in the

PORT VERTICAL GYRO mode. (A separate Model 1080Y

vertical gyroscope provides STARBOARD GYRO attitude

information for display on the pilot instrument panel, attitude

indicator, and pitch/roll to the ASE when in the STBD VERTI-

CAL GYRO mode.) The A/A24G-39 AHRS is the only

source of heading information for the ASE yaw channel and the

aircraft avionics systems (tacan, TACNAV, and cockpit BDHI,

RMI, and course indicator displays).

The source of the magnetic heading information used by

the AHRS is a remote compass transmitter unit located within

the aft tail cone area. Loss of reliable attitude and/or heading

from the AHRS is evidenced by illumination of the COMPASS

Figure 2-40. Pylon Folded

FAIL light on the caution panel, and display of the OFF flag on

the copilot attitude indicator. The standby compass located on

2.15 FLIGHT INSTRUMENTS

the cockpit instrument panel glareshield is a completely

separate backup wet compass provided for instances of

2.15.1 Standby Compass. A magnetic standby compass is

complete loss of AHRS heading information or aircraft

at the top center of the instrument panel. A standby compass

electrical power.

correction card is on the pilot side of the instrument panel.

Electrical power to the AHRS may be controlled only by

2.15.2 Free-Air Temperature Gauge. A bimetallic free-air

engaging or disengaging the circuit breakers. Power for the

temperature gauge is on the centerline of the helicopter on the

COMPASS FAIL caution panel light is supplied by the COMP

windshield glass panel.

dc breaker located in the pilot circuit breaker panel (Figure 2-

23). Power for the AHRS is supplied by the COMP ac breaker

2.15.3 Clocks. Two

8-day,

12-hour, elapsed time clocks

located in the copilot circuit breaker panel (Figure 2-23).

(Figure 2-7) are installed on the instrument panel. The control

knob for the elapsed time mechanism is at the upper right

Prior to application of power to the AHRS, adjust the two

corner of the clock face. The clock is stem wound and stem set

copilot attitude indicator trim knobs so that each knob arrow

with a knob in the lower left corner of the clock face.

points to the black reference dot (Figure

2-43). After initial

power application, a

2-minute warmup and gyro erection

2.15.4 Pitot-Static System. Two pitot tubes with static

period is required before reliable attitude information is

ports are mounted over the cockpit, one on the right side and

available.

(During this period, the COMPASS FAIL caution

the other on the left side of the helicopter, forward of the engine air

light and the copilot attitude indicator flag are visible.) Any re-

intakes. The lines carrying static pressure from both pitot-static

sidual attitude errors remaining after warmup can be eliminated

tubes are connected together so that the pilot and copilot

by depressing the ERECT button located on the AHRS control

airspeed and vertical velocity indicators, the TAS transducer,

panel (Figure 2-44) until the copilot attitude indicator displays

and BAR ALT controller are fed from a common static line.

satisfactory pitch and roll values. (If erected with the UH-3H

The starboard pitot pressure line feeds the pilot airspeed

on a flat and level ramp, the copilot attitude indicator will

indicator and the TAS transducer. The port pitot pressure line

display level roll attitude and a slight noseup attitude in pitch.)

feeds only the copilot airspeed indicator. A restrictor is

Attitude correction can also be performed during straight and

installed in the static pressure line to the BAR ALT controller

level flight by use of the ERECT button.

sensing unit to filter transient pressure changes, thereby

contributing to a more stable flight-path. The TAS transducer

Note

converts pitot-static airspeed information to an electronic TAS

signal that is fed to the TACNAV or navigation system.

· The quality of electrical power available at ship

Sleeves are installed on both pitot tubes to ensure correct

flight deck launch spots is so poor that AHRS

airspeed readings and to prevent erroneous inputs to the BAR

reliability can be severely degraded. COMPASS and

ALT controller when the ice shield is installed. Both pitot tube

PHASE B (ÆB) circuit breakers should be pulled

heads may be heated to prevent icing. Refer to Pitot Heaters,

when the UH-3H has deck edge power applied, and

paragraph 2.18.5.

engaged only with another source of external power

or after operating on helicopter power.

2-63

ORIGINAL

NAVAIR 01-230HLH-1

· When operating aboard ship, the deck

Note

motion in pitch, roll, and yaw will be reflected

in the AHRS outputs.

The SYNC meter needle does not function when

helicopter accelerations exceed preset values,

· Use of the ERECT button should be

such as during turns. In these flight conditions,

limited to a maximum period of 3 minutes to

the needle is stowed in the center of the meter

avoid thermal reliability problems. After a 1-

and no needle oscillations are seen. This stowed

minute wait, the ERECT button may be used

needle condition during maneuvers should not be

again.

mistaken for properly synchronized AHRS

heading operation.

· Because of the high inertia of the AHRS

Note

gyro, it is still spinning down for 20 minutes

after electrical power is removed, and is

vulnerable to reliability problems if the

During flight, the COMPASS FAIL advisory

helicopter is moved in an erratic fashion or if

light will occasionally illuminate momentarily as

electrical power is reapplied before rundown is

part of the normal AHRS operation. This will

usually occur upon rolling out of turns or upon

completed. To minimize AHRS gyro failures,

selecting the SLAVED mode if the DG or

avoid reapplying electrical power to the gyro

once 45 seconds have elapsed. Also, minimize

EMERG modes have been in use. Brief heading

the severity of helicopter towing impulses

fluctuations may also be seen on the cockpit

using steady and smooth starting and stopping

instruments during mode changes.

techniques.

The DG mode of operation is normally used when the

Earth's magnetic field is an unreliable reference (i.e., in polar

SLAVED is the UH-3H helicopter primary mode of AHRS

heading operation. In this mode, heading is furnished by a

regions or when the remote compass transmitter has failed).

directional gyro that is continually corrected for magnetic

Aircraft heading is furnished only by the directional gyro and is

heading by the remote compass transmitter (e.g., the gyro is

subject to error buildup if the correct latitude and hemisphere

slaved in the raw magnetic heading source). For this normal

are not inserted into the LAT and N/S controls located on the

control panel (Figure 2-44). The DG mode allows the pilot to

mode of operation, set the mode selection switch on the com-

set the AHRS heading to any selected reference by pushing and

pass control panel (Figure

2-44) to the SLAVED position for

the entire mission. Verify that the hemisphere switch is set to

turning the HDG PUSH knob until the desired heading is

the N position if in the northern hemisphere or S if operating in

displayed on the cockpit indicators and TACNAV. In polar

southern latitudes. Set the LAT control to the approximate

regions, the DG mode heading estimate should be based on the

average latitude expected for the mission. Hemisphere and

best available information to avoid large navigation errors.

latitude settings are especially important for minimizing

The EMERG mode of operation is used only when the

heading error buildup during sustained periods of helicopter

maneuvering (such as starboard delta patterns).

normal heading information from the AHRS is unreliable.

Magnetic information is passed directly from the remote

After AHRS warmup is completed and the COMPASS

compass transmitter for use until the AHRS can be replaced.

FAIL advisory light has extinguished, the directional gyro must

However, the magnetic heading information in the EMERG

mode is not as accurate and is not damped by the directional

be synchronized to the remote compass transmitter by

gyro. Since no automatic reversion to the mode occurs for

depressing the HDG PUSH knob until the SYNC meter needle

is centered (Figure 2-44). After takeoff (when at altitude with

AHRS gyro failures (COMPASS FAIL light or copilot altitude

a straight and level attitude), check the SYNC meter again to

indicator flag visible), the EMERG mode must be selected

determine whether initial magnetic heading errors because of

manually.

ship or helicopter launch pad magnetic fields have biased the

AHRS initial synchronization. If the needle is not centered,

push the HDG PUSH knob to remove the errors. To avoid

TACNAV navigation error buildup, the SYNC meter needle

should be checked periodically.

Because of the oscillatory nature of the compass

information in the EMERG mode, the pilot

should immediately discontinue the use of the

ASE yaw channel.

2-64

ORIGINAL

NAVAIR 01-230HLH-1

Note

2.15.7 Attitude Indicators. Two attitude indicators (Figure

2-43) on the instrument panel visually indicate the helicopter

If it is necessary to troubleshoot the system when

attitude. The indicator face consists of a stationary miniature

in the EMERG mode with the yaw channel

aircraft representing the helicopter, a bank angle scale, bank

engaged, follow these procedures to avoid

index, and a moving two-colored sphere with a distinct white

driving rudder pedal forces when returning to the

horizontal line dividing the two colors, white above and black

SLAVED mode. Place feet on rudder pedals and

below. A warning flag marked OFF will appear in the face of

disengage the yaw channel. Switch to the

the indicator when the indicator is inoperative. The warning

SLAVED mode. Synchronize the compass

flag will appear until about 68 seconds after ac power has been

heading by pressing the HDG PUSH knob until

applied to the circuit, when any unbalance of the three phases

the SYNC needle is centered. Cheek that the

of ac power occurs, or the directional gyro system for the

BDHI/ RMI compass card and standby compass

indicator has failed. Two trim adjustment knobs are on the

are within ±5^o of each other (if not, repeat the

front of the attitude indicators, one at the lower left of the panel

procedure until they are). Engage the yaw

for adjusting roll, and the other at the lower right of the panel

channel and remove feet from rudder pedals.

for adjusting pitch. The pilot attitude indicator operates on ac

power and is protected by a circuit breaker. The copilot attitude

2.15.6 Compass System Control Panel. This panel

indicator operates on ac power and is protected by circuit

(Figure 2-44) is on the center console. When in the SLAVED

breakers.

mode, pressing and holding down the HDG PUSH knob

automatically causes synchronization of heading outputs of the

Each attitude indicator also has a turn-and-slip indicator

directional gyro with the compass transmitter heading. In the

mounted on the bottom that visually indicates the helicopter

DG mode, the HDG PUSH knob may be pushed and turned

rate of turn and flight condition.

clockwise or counterclockwise to set the heading on the BDHIs

and RMI, and the system works with the directional gyro.

Note

In the SLAVED mode, the SYNC window indicates

The warning flag marked OFF on the copilot gyro

synchronization of the directional gyro in the compass

will appear momentarily when the helicopter is in a

transmitter heading. The EMERG position provides emergency

high vibratory flight regime. This may be when the

operation using only the compass transmitter and is used when

helicopter is entering a hover or in any situation of

a directional gyro failure has occurred.

impending translational lift. The warning flag will

disappear when the helicopter leaves the high

The N/S switch selects polarity of latitude correction

vibratory flight regime.

signal: N for northern hemisphere and S for southern

hemisphere. The LAT knob selects latitude of flight (0^o to 90^o)

2.15.8 Turn Rate Switches. Two turn rate switches (Figure

to correct for heading gyro drift because of Earth rotation. The

2-7), one each for the pilot and copilot, are on the instrument

N/S switch and LAT knob must be set for the SLAVED mode

panel. The switches are marked TURN RATE and have marked

as well as DG mode since the system automatically switches

positions NORM and ALT. When the switches are placed to

from SLAVED to DG mode whenever the helicopter is in a

NORM, the copilot indicator receives rate of turn information

turn.

from the ASE rate gyro amplifier demodulator, and the pilot

indicator receives rate of turn information from the dc powered

The ERECT pushbutton, when pressed, increases the roll

rate gyro transmitter. This system remains operative when the

erection rate of the vertical gyro.

OFF flag in the attitude indicator appears, provided dc power is

available. When either switch is placed to ALT, that indicator

receives rate of turn information from the other system.

Figure 2-44. Compass System Control Panel

2-65

ORIGINAL

NAVAIR 01-230HLH-1

Figure 2-43. Attitude Indicator

2.15.9 Altitude Encoder. The altitude/encoder (Figure

foot change in altitude and drives a drum that also

2-45) functions as a barometric altimeter for the pilot and

indicates altitude in hundreds of feet. The

100-foot drum

a pressure altitude sensor for the AN/APX-72 IFF

will cause the 1,000-foot drum to index one digit when it

transponder.

moves through the short period between 9 and 0, and the

1,000-foot drum actuates the 10,000-foot drum in a similar

manner. During the long period between

0 and 9, the

1,000- and 10,000-foot drums are locked in position. In the

space corresponding to zero, the

10,000-foot drum is

marked with striped lines. The pointer acts as a vernier of

the 100's drum as well as being an indication of trend

information. The barometric set knob permits altimeter

setting from 28.10 to 31.00 inches Hg and these settings

will be displayed by the barometric pressure setting

window. The combined readings of the drums and pointer

indicate an altitude measurement in thousands and

hundreds of feet based on the selected altimeter setting.

Two techniques may be used to read indicated altitude

on the drum and pointer-type altimeter. Read the drums

without referring to the 100-foot pointer as a direct digital

readout of both thousands and hundreds of feet, or read the

digital readout in thousands without referring to the 100-

foot drum, and then add the 100-foot pointer indication.

Figure 2-45. Altitude Encoder

The 100-foot pointer serves as a precise readout of values

less than 100 feet required for determining lead points for

2.15.9.1 Altimeter. The AAU-21/A altitude indicator

level-off altitudes, maintaining level flight, and during

(Figure 2-7) is on the pilot side of the instrument panel.

instrument approaches. The system is equipped with a

The altimeter operating range is from -1,000 to +38,000

continuously operating vibrator to improve altitude

feet. The face of the instrument indicates readings in feet

measuring accuracy. Power to operate the vibrator is

in 50-foot units with the numerals representing hundreds

furnished through a dc circuit breaker marked IFF TEST.

of feet. The pointer makes one revolution for every 1,000-

2-66

ORIGINAL

NAVAIR 01-230HLH-1

2.15.9.2 Encoder. The AAU-21/A encoder provides a

warning index marker on the indicator. Depressing the

digital output of pressure altitude in units of 100 feet to the

PUSH-TO-TEST control knob provides a testing feature of

AN/APX

72 IFF transponder for automatic pressure

the system at any time and altitude, provided the CPLR

altitude transmission. The encoder's operating range is

mode of ASE is disengaged. When the PUSH-TO-TEST

from - 1,000 to +38,000 feet and has a permanent altimeter

control knob is depressed, a visual indication of 100

±15

setting of 29.92. Loss of 115-vac, 400-Hz power will cause

feet on the indicator indicates satisfactory system

the warning flag marked CODE OFF on the pilot altimeter

operation. Releasing the PUSH-TO-TEST control knob

indicator to be displayed. Power to operate the encoder is

restores the system to normal operation. A low-level

furnished through an ac circuit breaker marked IFF IND.

warning light on the lower right corner of the indicator will

light and show the marking LOW anytime the helicopter is

Note

at or below the low-altitude limit that has been selected.

Loss of system power or tracking condition will be

Only when the altimeter is set to 29.92 will

indicated by a black and yellow striped flag that appears in

the transmitted altitude be the same as the

the indicator window on the lower center portion of the

displayed altitude.

indicator. If the system should become unreliable, the black

and yellow striped flag will appear and the indicator

2.15.9.3 Altimeter

(AAU-24/A). The altimeter (Figure

pointer will go behind a mask marked NO TRACK to

2-7), mounted on the copilot side of the instrument panel,

prevent erroneous readings. While hovering at low

is identical to and operates in the same manner as the pilot

altitudes over smooth surfaces, fluctuations of about ±1.5

AAU-21/A altitude encoder except that there is no warning

feet can be expected for altitudes under 100 feet. During

flag on the indicator.

normal flight operations above

1,000 feet, the altimeter

indicates an unreliable condition.

2.15.10 Radar Altimeter

The system requires both ac and dc power for

operation and is protected by a circuit breaker.

After turning equipment on, allow about 3

minutes for systems to reach operating

temperature. To avoid damage to system

components, do not place system in

operation until 3 minutes have elapsed since

system was last turned OFF.

The radar altimeter system

(AN/APN-171(V))

provides instantaneous indication of actual clearance

between the helicopter and terrain from 0 to 1,000 feet.

Altitude in feet is indicated by the two radar altimeter

indicators on the instrument panel in front of the pilot and

copilot. The radar altimeter

(Figure

2-46) contains a

Figure 2-46. Radar Altimeter

pointer that indicates altitude on a linear scale from 0 to

100 feet in

2-foot units, 100 to 200 in 10-foot units, and

2.15.11 Radar Altitude Warning System. The

200 to 1,000 in 50-foot units. A control knob on the lower

AN/APQ-107 RAWS provides the pilot and copilot with

left corner of the indicator combines functions to serve as a

aural and/or visual warning signals of low-altitude con-

test switch, a low-level warning index set control, and an

ditions and/or failure of the radar altimeter. The visual

on/off power switch. The system is turned on by turning

indication is provided by a pulsating ALTITUDE light on

the control knob marked PUSH-TO-TEST clockwise from

the caution panel. The aural tone is channeled through the

OFF, and is the only control necessary for equipment

intercommunications system to pilot and copilot headsets.

operation. Operation of the control knob on either indicator

These signals are triggered under any of the following

will turn the system on; however, both control knobs must

conditions:

(1) Whenever the radar altimeter decreases

be turned fully counterclockwise to turn the system OFF.

Continued clockwise turning of the control knob toward

through 100

±5 feet when the coupler is not engaged or

the SET position will permit the pilot to select any desired

reaches 30 ±5 feet with the landing gear up, (2) When the

low-altitude limit that will be indicated by the low-level

radar altimeter output signals are unreliable and if

2-67

ORIGINAL

NAVAIR 01-230HLH-1

Figure 2-47. RAWS Warning - All Conditions

electrical power to the radar altimeter fails (Figure

2-47).

circuit, to indicate a particular system is in operation or an

A test switch marked RAWS TEST is on the instrument

unsafe condition exists. When a system is in operation or

panel to the right of the caution panel. To test system

an unsafe flight condition exists, the advisory light for that

reliability, make certain that the radar altimeter is

particular system or condition comes on and remains on

energized and ac and dc power is being supplied, and then

until the system is turned off or the unsafe flight condition

depress the test button. The aural and visual warning

is corrected. Pressing the switch marked TEST on the

signals should occur simultaneously. The system operates

caution panel tests lights of both the caution and advisory

115-vac power and

28-vdc power. The system is

panel.

protected by circuit breakers.

2.16.1.1 Advisory Panel Lights. Refer to Figure 2-48

for a detailed description of the advisory panel.

WARNING

2.16.2 Caution Panel. The caution panel (Figure

49) marked CAUTION PANEL is on the pilot side of the

instrument panel. The caution panel gives the pilot visual

With the landing gear down, the RAWS will

indication of failure or unsafe conditions of certain critical

not provide an indication of an unreliable

power equipment in the helicopter. The caution panel

radar altimeter.

contains placard-type amber caution lights, each having its

own operating circuit, to indicate a particular condition in

2.16 WARNING, CAUTION, AND ADVISORY

the helicopter. If a failure or unsafe condition occurs in one

LIGHTS

of the systems, the caution light for that particular

condition remains on until the failure or unsafe condition is

2.16.1 Advisory Panel. The advisory panel (Figure 2-

corrected. A master light marked MASTER CAUTION

48) marked ADVISORY PANEL is on the copilot side of

PUSH TO RESET is on either side of the instrument panel.

the instrument panel. The advisory panel gives the pilots

The master light goes on when any of the caution panel

visual indication of certain operating conditions in flight or

lights are energized by a malfunction. The master light will

while on the ground. The advisory panel contains placard-

type green advisory lights, each having its own operating

2-68

ORIGINAL

NAVAIR 01-230HLH-1

remain on until the malfunction is corrected or until

panel. The circuit breaker marked PWR provides electrical

deenergized by the pilot. Pressing the light deenergizes the

power for the normal operation of the caution lights and

master light, permitting the master light to indicate a

the circuit breaker marked TEST provides power for the

second malfunction if one should occur while the first

test circuit only.

malfunction is still present. The MASTER CAUTION

lights are simultaneously dimmed when the caution panel

2.16.2.1 Caution Panel Lights. Refer to Figure 2-49

is dimmed. This is done using the PILOTS FLIGHT INST

for a detailed description of the caution panel.

LTS rheostat on the overhead switch panel. The caution

lights operate through circuit breakers marked WARNING

LTS PWR and TEST that are on the center circuit breaker

EXT PWR ON

#1 ENG ANTI-ICE ON

#2 ENG ANTI-ICE ON

PARKING BRAKE ON

NOSE DOOR OPEN

CHECK BLADE FOLD

01771122

ADVISORY LIGHTS

INDICATION

CONDITION

ACTION

EXT PWR ON

AC power receptacle door

Light will extinguish when

is open.

door is secured.

# 1 ENG ANTI-ICE ON

No. 1 anti-ice switch is on.

Light will extinguish when

# 2 ENG ANTI-ICE ON

No. 2 anti-ice switch is on.

engine anti-ice switch is

placed off.

PARKING BRAKE ON

Parking brake handle not

Reset and release parking

fully retracted.

brake.

NOSE DOOR OPEN

Electronic compartment

Check door security.

door is open or not fully

secure.

CHECK BLADE FOLD

DC power is applied and

If blade spread, call trouble -

blade fold interlock

shooter.

system is not in flight

condition.

Figure 2-48. (NON-ET) Advisory Light Panel

2-69

ORIGINAL

NAVAIR 01-230HLH-1

CAUTION PANEL

#1 GENERATOR

PRI SERVO PRESS

#2 GENERATOR

#1 RECTIFIER

AUX SERVO PRESS

#2 RECTIFIER

FWD FUEL LOW

TRANS OIL PRESS

AFT FUEL LOW

LAMP

TEST

FWD FUEL BYPASS

TRANS OIL HOT

AFT FUEL BYPASS

#1 INLET ANTI-ICE

ALTITUDE

#2 INLET ANTI-ICE

TAIL TAKE OFF

MAIN TRANS CHIP

ROTOR BRAKE ON

INTMED TRANS

IFF

TAIL TRANS CHIP

FUEL DUMP

COMPASS FAIL

BLADE PRESS

CAUTION LIGHT

CAUSE/ DESCRIPTION

ACTION*

#1 GENERATOR

Associated generator contactor relay

Recycle generator; if recycle fails, -

#2 GENERATOR

opens and falls off line.

LAND AS SOON AS PRACTICAL.

If both fail, - LAND AS SOON AS

POSSIBLE.

#1 RECTIFIER

Failure of rectifier or associated reverse

Single rectifier - LAND AS SOON

#2 RECTIFIER

current cutout relay.

AS PRACTICAL. Both rectifiers -

LAND AS SOON AS POSSIBLE.

PRI SERVO PRESS

Pressure falls below 1,000 psi or 1,000

LAND AS SOON AS POSSIBLE. If

AUX SERVO PRESS

psi switch failure.

pressure is below normal, secure

affected system.

FWD FUEL LOW

210 to 280 pounds per tank 3⁰nose

Avoid altitudes above 6⁰nose up,

AFT FUEL LOW

down - 170 to 200 pounds per tank in a

crossfeed, land as fuel remaining

hover.

permits.

FWD FUEL BYPASS

Pressure drop of 1.1 to 1.7 psi at the

Crossfeed from the good tank. -

AFT FUEL BYPASS

fuel filter.

LAND AS SOON AS PRACTICAL.

TRANS OIL PRESS

Main gearbox oil pressure is below 3.5

LAND AS SOON AS PRACTICAL.

psi at the forward right corner of the

Check for other indications.

main gearbox.

TRANS OIL HOT

Main gearbox oil temperature above

LAND AS SOON AS PRACTICAL.

120⁰ at oil cooler outlet.

Check for other indications.

* Procedures contained in the action column of this matrix shall not supersede the more detailed EMERGENCY

PROCEDURES of chapter 12.

Figure 2-49. Caution Light Panel (Typical) (Sheet 1 of 2)

2-71

ORIGINAL

NAVAIR 01-230HLH-1

CAUTION LIGHT

CAUSE/DESCRIPTION

ACTION*

MAIN TRANS CHIP

Metallic particles in main gearbox oil.

LAND AS SOON AS PRACTICAL.

Check for other indications.

INTMED TRANS CHIP

Metallic particles or over heat condition

LAND AS SOON AS POSSIBLE.

TAIL TRANS CHIP

in the associated gearbox oil.

Check for other indications.

#1 INLET ANTI-ICE

Anti-ice system failure or inability to

Check circuit breaker; continue flight

#2 INLET ANTI-ICE

maintain 38⁰inlet temperature.

as conditions permit.

ALTITUDE

Activation of the RAWS.

Check altitude.

TAIL TAKE OFF

Tail takeoff free wheel unit

Check N_r. -LAND AS SOON AS

failure.

POSSIBLE. (Except for low N_r).

Tail takeoff warning system

failure.

Nr at 96 percent or below.

ROTOR BRAKE ON

Manual or automatic rotor brake

On deck - secure both speed selectors.

pressurized.

In flight - LAND AS SOON AS

POSSIBLE.

IFF

Mode 4 code zeroized.

MASTER SWITCH to

Transponder is not replying to

NORM.

Mode 4 interrogations.

Check Mode 4 toggle is on.

COMPASS FAIL

A/24g compass has failed.

Switch to STBD VERTICAL GYRO

Synchro interface unit (SIU)

on CHANNEL MONITOR panel to

has failed.

preclude erroneous ASE inputs.

VERTICAL GYRO in the fast

erect mode.

FUEL DUMP

Fuel dump valve is open.

Monitor fuel remaining.

1. Airspeed - 80 KIAS, minimize

BLADE PRESS

Loss of blade pressure.

maneuvering.

2. Altitude - Minimum safe.

3. Blade Pressure Circuit Breaker -

Check in.

4. LAND AS SOON AS

PRACTICABLE

* Procedures contained in the action column of this matrix shall not supersede the more detailed EMERGENCY

PROCEDURES of chapter 12.

Figure 2-49. Caution Light Panel (Typical) (Sheet 2 of 2)

2-72

ORIGINAL

NAVAIR 01-230HLH-1

CAUTION PANEL

#1 GENERATOR

PRI SERVO PRESS

#2 GENERATOR

LAMP

#1 RECTIFIER

AUX SERVO PRESS

#2 RECTIFIER

TEST

FWD FUEL LOW

TRANS OIL PRESS

AFT FUEL LOW

FWD FUEL BYPASS

TRANS OIL HOT

AFT FUEL BYPASS

#1 INLET ANTI-ICE

ALTITUDE

#2 INLET ANTI-ICE

TAIL TAKE OFF

MAIN TRANS CHIP

ROTOR BRAKE ON

FWD DOOR OPEN

INTMED TRANS

IFF

AFT DOOR OPEN

TAIL TRANS CHIP

FUEL DUMP

COMPASS FAIL

BLADE PRESS

CAUTION LIGHT

CAUSE/DESCRIPTION

ACTION

FWD DOOR OPEN

Forward door is open or

Check door security.

not fully secure.

AFT DOOR OPEN

Aft door is open or not

Check door security.

fully secure.

Figure 2-49.1. (ET) Caution Light Panel

2.17 FIRE DETECTOR SYSTEM

2-9). A light on the instrument panel and a light in either

the No. 1 or No. 2 engine fire emergency shutoff selector

Two fire detector systems (one for each engine) are

handle will go on in case of a fire in the corresponding

installed to warn the pilot of an engine fire. Three

engine compartment. To test the engine fire detector

hermetically sealed temperature sensitive sensing loops

system, place the spring-loaded switch to the up FIRE

and two interconnecting cables that have no heat sensing

TEST position. The fire warning lights on the instrument

function are in each engine compartment. They are wired

panel and in the fire emergency shutoff selector handle

into a closed series loop connected to a control unit that

will go on. The switch will return to OFF when released,

turns on warning lights in the pilot compartment in case

and the lights will go off.

of a fire. The engine fire detector systems operate on ac

power. The Nos. 1 and 2 fire detector systems receive

2.18 HEATING SYSTEM

power from the inverter when the generators are

inoperative.

The heating system (Figure 2-50) consists of a fan, an

internal combustion heater, a plenum chamber, and ducts

2.17.1 Fire Warning Lights and Test Switch. Two

that run along the right side of the cabin wall and into the

red engine fire warning lights and test switch (see Figure

pilot compartment. A fire access port is in the aft heater

2-7) are on a plate marked FIRE WARN on the pilot side

duct aft of the broom closet on the starboard bulkhead. In

of the instrument panel. The lights are marked NO. 1

addition, a spring-loaded fire extinguisher door is in the

ENG and NO. 2 ENG. The switch has two marked

heater compartment access cover on the aft side of the

positions FIRE TEST and OFF. In addition, four red

ASE compartment. The heater unit in the heater

engine fire warning lights, two for each engine, are in the

compartment next to the control enclosure in the cabin

engine fire emergency shutoff selector handles marked

operates on fuel pumped from the forward fuel tank by a

FIRE EMER SHUTOFF SELECTOR NO.

1 ENGINE

heater fuel pump cycling valve to the heater unit, where it

and NO. 2 ENGINE on the overhead switch panel (Figure

is ignited by a spark plug. Fuel consumption of the heater

2-73

ORIGINAL

NAVAIR 01-230HLH-1

unit, operating continuously in the HIGH position, is 1.2

2.18.3 Heating and Ventilating Diffusers. A heater

gallons

(8 pounds) per hour. The spark plug operates

diffuser and register are in each heater duct that extends

electrically on

28-vdc current from the monitored bus,

along the outside of the pilot compartment, just above the

boosted by the heater ignition unit mounted in the heater

floor. Four diffusers are in the heater duct that extends

compartment. Air is drawn into the heater intake port on

along the right side of the cabin on the floor. Knobs

the fuselage above and to the rear of the pilot window,

marked OPEN and CLOSED are used to regulate the flow

and then through a heat exchange unit surrounding the

of warm air through the diffusers. The diffusers and fan

combustion unit. Heated air is then blown into the plenum

are also used as a ventilating system.

chamber, the cabin heater duct containing four diffusers,

and the pilot compartment heater ducts, each containing

2.18.4 Normal Operation.

two diffusers. The fan also supplies air to the heater

combustion chamber.

1. Heater switch - HIGH OR LOW.

2.18.1 Cabin Heater Switch. The heating system is

operated by a switch, marked CABIN HEATER with

three marked positions LOW, OFF, and HIGH on the

WARNING

overhead switch panel (Figure

2-9). The heater switch

controls the heater fuel pump and cycling valve and the

ignition unit. When the switch is at LOW, the heater will

Hot exhaust from the heater may burn

maintain a temperature of about 65^oC in the ducts. When

personnel in the vicinity of the heater

the switch is at HIGH, the heater will automatically

exhaust.

maintain a temperature of about 149 ^oC in the ducts. An

overheat switch will shut off the heater if for any reason

2. Heater diffusers- ADJUST AS DESIRED.

the heat in the plenum chamber rises to 177^oC. During

the start sequence of the heater, if the fuel does not ignite

within 40

±5 seconds, the thermal switch in the exhaust

tube will not actuate, causing the overheat relay circuit to

be completed through the contacts of the time delay relay.

These contacts close after 40

±5 seconds, shutting off

electrical power and fuel flow to the heater. The heater

Before operating heater, check all heater

may be restarted by recycling the HEATER HEAT circuit

outlets for clearance from luggage,

breaker

(CAB HTR UH-3H) with the heater control

clothing, inflatable equipment, and other

switch OFF. The heater will also shut off if the fan fails to

gear that might be damaged by heat.

operate. The heater switch is energized by the 28-vdc

monitored bus and is protected by a circuit breaker.

Note

2.18.2 Cabin Heater Fan Switch. The switch marked

The heater is shut off by moving the heater

FAN and mounted on the overhead switch panel (Figure

switch OFF. The fan will continue to

2-9) has two positions marked OFF and ON. The fan

operate after the heater is shut off until the

switch controls a relay connecting 115-vac power to the

temperature in the plenum chamber drops

heater fan (blower) in the upper part of the heater unit.

49^oC. If the temperature in the plenum

The fan switch is energized by the 28-vdc monitored bus

chamber should rise to 49^oC during hot

and is protected by a circuit breaker. Placing the cabin

weather, the fan will begin to operate

heater fan switch ON without operating the cabin heater

whenever the monitored bus is actuated.

switch will draw outside air into the heater system and

ventilate the pilot compartment and cabin.

2.18.5 Pitot Heaters. A pitot heater switch marked

PITOT HEAT with marked position ON is on the over-

head switch panel (Figure

2-9). When placed ON, an

electric heater in each pitot head is turned on to prevent

ice formation in the pitot head. Each pitot heater operates

on dc power and is protected by a circuit breaker marked

PITOT HEAT on the center circuit breaker panel.

2-74

ORIGINAL

NAVAIR 01-230HLH-1

2.18.6

(ET) ENVIRONMENTAL AIR SYSTEM

system includes the heater and basic controls for

The environmental air system provides for the

manual or thermostatically controlled operation as

heating, cooling, or ventilating of the cockpit and

well as associated fuel and air systems. The air

cabin (Figure

2-50.1).

52,000 BTU/hour of cooling

conditioning system includes the air conditioner

is provided at the evaporator outlet throughout the

comp onents and basic controls to thermostatically

aircraft flight profile and during ground operation.

cool or ventilate the cockpit or cabin. The air

The system may also be operated prior to engine start

conditioning system requires 115/200 volt, 3 phase,

through the use of the sponson-mounted auxiliary

400 Hz alternating current for operation. Electrical

power unit. The system is divided into functional

power for the control circuitry is supplied by the

sections of temperature control, distribution, heating,

monitored bus at 28 volts dc through the HEATER

and air-conditioning. The temperature control system

HEAT and CABIN VENT circuit breakers on the No.

includes the components that automatically control

1 junction box. The No. 1 generator, at 200 volts ac

heater or air-conditioner operation to maintain the

through the AIR COND CONDENSER circuit

desired temperature. The monitored dc bus supplies

breaker, supplies electrical power for the condenser

electrical power for control circuitry.

The

motor. Electrical power for the compressor motor,

distribution system directs heated air to the cockpit

air conditioner vent blower, and the heater vent

and cabin diffusers, and cooled or ventilated air to

blower is supplied by the No. 1 generator through the

both diffusers and gaspers in the cockpit and cabin.

AIR COND COMPR MOTOR, HEATER BLOWER,

The heating system includes the heater and basic

and the AIR COND VENT BLO circuit breakers.

controls for manual or thermostatically controlled

The air conditioning system provides cool air to the

operation as well as associated fuel and air systems.

cockpit and cabin compartments. The system is

In manual operation, the heater is operated

charged with HFC 134a that is readily vaporized or

continuously; while during automatic operation, the

changed back to a liquid at low temperatures and

cabin heat level and air-conditioning heat-loss level is

pressures.

When the system is operating, the

controlled

by setting

the

CABIN AIR

refrigerant is compressed, which reduces the overall

TEMPERATURE rheostat at the crew chief’s ICS

volume. The heat of compression is given up to the

station. The air-conditioning system includes air-

outside air as the high-pressure vapor passes through

conditioner components and basic controls to

the condenser, which is cooled by the condenser

thermostatically cool the cockpit and cabin. The

blower. This heat transfer changes the vapor into a

ventilation system utilizes the distribution system and

high-pressure liquid, which is collected in the filter

the heater or air conditioner blower to provide

drier. The liquid is subjected to a condition of

ventilated air to the diffusers and gaspers in the

reduced pressure and increased surface area upon

cockpit and cabin.

passing through the expansion valve into the coils of

the evaporator. The liquid vaporizes in the coils.

2.18.6.1 (ET) Air Conditioning System. The

The heat necessary for vaporization is taken from the

air conditioning system equipment is divided into

air that is to be cooled. The vaporized liquid is then

functional

sections

of temperature control,

passed back into the motor-compressor, where the

distribution, heating, and air conditioning.

The

process begins again.

temperature control system consists of an AIR

COND CONTROL panel, a CABIN AIR

2.18.6.2

(ET) Manual Operation. When all

TEMPERATURE rheostat, sensing elements, and a

heater/air conditioner system circuit breakers are

control unit.

These units provide an automatic

engaged and the SELECTOR switch is positioned to

thermostatically controlled temperature in the cockpit

HEATER, electrical power energizes the heater

and cabin. The heating system can be operated

blower relay and air conditioner blower relay and the

manually or automatically, and the blowers may be

blowers start. When the heater blower has delivered

operated without the heater or air conditioner system

sufficient air to the heater combustion chamber, the

to ventilate the cockpit or cabin. The air conditioner

air pressure switch closes, directing electrical power

system is operated automatically to cool the cockpit

through the overheat relay to operate the master fuel

and cabin. In manual operation, the heater is

valve and the ignition unit. When the HEAT MODE

operated continuously, while during automatic

switch is positioned to MAN OVRD, a circuit

operation, setting the CABIN AIR TEMPERATURE

through the 140.6°C (285°F) cycling thermal switch

rheostat on the control panel, at the crew chief’s ICS

and the HEAT MODE switch opens the fuel control

station, controls the cabin heat level and the air

unit solenoid valve, passing fuel to the heater. When

conditioning heat-lost level. The distribution system

the ignition unit and the fuel valve are energized, the

directs heated, cooled, or ventilated air to the

45-second time delay relay is simultaneously

diffusers in the cockpit and cabin. The heating

energized through the 93.3°C (200°F) exhaust

2-76

ORIGINAL

NAVAIR 01-230HLH-1

thermal switch. If exhaust temperature does not reach

drops.

The sensing elements or the temperature

93.3°C (200°F) within

45 seconds after starting the

selector then unbalance the heater control unit bridge to

heater, the time delay relay closes and energizes the

start the heater. If the heater control unit should

176.7°C (350°F) overheat relay, which interrupts power

malfunction when the heater system is operating, the

to the fuel valve and ignition unit and shuts off the

140.6°C (285°F) thermal switch will cycle the heater as

heater. Cycling the SELECTOR switch on the control

during manual operation. The thermal switch will start

panel may then restart the heater. With the air fuel

the heater when the temperature falls below 140.6°C

mixture and ignition spark being supplied, the heater

(285°F), and shut the heater off when the temperature

operates, supplying heated ventilating air for the

exceeds

140.6°C

(285°F).

If the plenum chamber

cockpit and cabin and discharging exhaust gases

approaches an unsafe temperature, the 176.7°C (350°F)

overboard. The heater continues to operate until the

normally closed thermal switch will open and de-

140.6°C (285°F) normally closed thermal switch opens

energize the overheat relay circuit to stop fuel and

and de-energizes the fuel control unit solenoid valve to

ignition. The blower will continue to operate to provide

shut off fuel supply. The blower continues to operate

safe cooling of the heater and proper exhaust of

until the plenum chamber temperature drops, and the

combustion gases.

140.6°C (285°F) thermal switch closes to resume heater

operation. If air in the plenum chamber approaches

2.18.6.4 (ET) Air Conditioning Operation. When

unsafe temperatures, the

176.7°C

(350°F) normally

all air conditioning circuit breakers are engaged and the

open thermal switch closes and energizes the overheat

SELECTOR switch is positioned to AIR COND,

relay. The overheat relay opens the heating system

electrical power energizes the air conditioning blower

circuit to stop fuel flow and ignition to the heater, while

relay to start the blower.

Electrical power

the blower continues to operate to bring the plenum

simultaneously flows to the sensing elements, the

chamber temperature down to safe limits. Under these

control unit

bridge and the CABIN AIR

circumstances, or when the HEAT MODE switch is

TEMPERATURE rheostat. The compressor motor,

OFF, the blowers, powered through the normally open

which is an integral part of the motor-compressor, and

48.9°C (120°F) thermal switch, will continue operating

the condenser motor, which is an integral part of the

regardless of the position of the selector switch. The

condenser, are started by action of their respective

thermal switch circuit bypasses the VENT AIR and the

relays when the SELECTOR switch, on the AIR COND

SELECTOR switches to continue blower operation

CONTROL panel in the cockpit, is positioned to AIR

until the plenum chamber temperature is below 48.9°C

COND. A COMPR light, on the crewman’s radio

control panel, illuminates while the motor-compressor

(120°F), to provide safe cooling of the heater and

proper exhaust of combustion gases.

The air

is running and goes out when the motor compressor

conditioner blower operates whenever the heater blower

stops. The motor-compressor and the condenser motor

is operating, to provide ventilating air in the upper ducts

relay energizing circuits are interlocked with the engine

which is controlled by gaspers in the ducts.

anti-ice system to prevent simultaneous operation of the

engine anti-ice system and the air conditioning system.

The air conditioning system supplies cooled air to the

2.18.6.3 (ET) Automatic Operation. When all

heater/air conditioner system circuit breakers are

cabin until the compartment sensing elements or the

engaged and the SELECTOR switch s at HEATER,

CABIN AIR TEMPERATURE rheostat balances the

power energizes the heater blower relay and air

control unit bridge. Once the bridge is balanced, the

conditioning blower relay and the blowers start. When

motor-compressor relay de-energizes and shuts off the

sufficient air is being delivered to the heater

motor-compressor. The motor-compressor also shuts

off when the pressure switch senses the pressure is

combustion chamber, the air pressure switch closes.

This causes power to be supplied through the overheat

below 10 psi or above 270 psi. A time delay relay in

relay to operate the master fuel valve and ignition unit.

the system prevents the compressor from coming on

When the HEAT MODE switch is positioned to AUTO,

until 15 seconds have elapsed since compressor shut-

off. The system is protected against extremely high

a circuit through the

140.6°C (285°F) cycling thermal

pressure by a relief valve set at

320 psi. When

switch and the HEAT MODE switch opens the fuel

maximum cooling of the cabin is desired, the CABIN

control unit solenoid valve, passing fuel to the heater.

AIR TEMPERATURE rheostat is turned to the

The air-fuel mixture and ignition spark assures heater

RECIRCULATE position and electrical power flows to

operation, thus supplying heated air for the cockpit and

the damper actuator relay, energizing the damper

cabin. The heater continues to operate until the sensing

actuator and closing the outside air duct. This allows

elements or the rheostat balances the heater control unit

the cabin air to recirculate through the air conditioning

bridge to shut the heater off. The blower continues to

operate until the plenum chamber or cabin temperature

2-77

ORIGINAL

NAVAIR 01-230HLH-1

system, cooling the air quickly.

Positioning the

thermistors, a fan, and a radio noise filter. One

SELECTOR switch to OFF will shut off the air

thermistor in each cabin air-sensing element is not used.

conditioning system.

The outside air temperature sensing element has one

thermistor and the air conditioning discharge sensing

2.18.6.5 (ET) Air Conditioner Control Panel. The

element has two thermistors. The heater discharge-

air conditioner control panel (Figure

2-50.2) contains

sensing element has two thermistors. One of the heater

the necessary switches and relays to control heating, air

discharge sensing thermistors is connected in series

conditioning, or ventilating operations.

The

with the outside air temperature sensing element, the

SELECTOR and the HEAT MODE switches control

cabin sensing elements, and one of the air conditioning

manual or automatic cycling of the heater. When the

discharge thermistors, through the control unit control

SELECTOR switch is positioned to HEATER,

relay. The other heater discharge sensing thermistor is

electrical power flows to the HEAT MODE switch,

connected in series with the CABIN AIR

sensing elements, control unit, CABIN AIR

TEMPERATURE rheostat through the control unit

TEMPERATURE rheostat, heater master fuel valve,

control relay and is also connected in series with the

and ignition unit. At this time, the heater may be

compartment sensing thermistor. The resistance in the

operated manually by positioning the HEAT MODE

sensing elements will vary to unbalance the control unit

switch to MAN OVRD, or automatically by positioning

bridge to start the heating or air conditioning system.

the HEAT MODE switch to AUTO. With the HEAT

MODE switch in the manual position, the heater will be

2.18.6.8

(ET) Heater Control Unit. The heater

cycled by the 140.6°C (285°F) thermal switch. With

control unit is located in the electronics compartment.

the HEAT MODE switch positioned to AUTO, the

It operates in conjunction with the CABIN AIR

heater will be cycled by the sensing elements, the

TEMPERATURE rheostat, AIR COND CONTROL

control unit, and the CABIN AIR TEMPERATURE

panel, and sensing elements to provide thermostatic

rheostat. The air conditioning system will start when

control of the cockpit and cabin temperature. The

the selector switch is positioned to AIR COND. The

control unit incorporates a bridge circuit, polarized

VENT AIR switch controls operation of the heater

switch, control relay, heat relay, and cooling relay. The

blower and air conditioning blower. When the VENT

temperature selector and the sensing elements vary the

AIR switch is positioned to ON, both blowers will

resistance to energize or de-energize the fuel control

circulate air through the cabin. The heater relay and the

solenoid valve. With the heating system operating, the

damper relay are also in the control panel. The

CABIN AIR TEMPERATURE rheostat or the sensing

SELECTOR switch actuates the heater relay, which

elements unbalance the bridge and actuate the polarized

energizes the heater blower relay and air conditioning

switch. At this time, the control relay is de-energized,

blower relay to start the blowers. The damper relay is

permitting current to flow through the CABIN AIR

energized to close the damper when the CABIN AIR

TEMPERATURE rheostat, the heater discharge sensing

TEMPERATURE rheostat is in RECIRCULATE

element, the outside air temperature sensing element,

the cabin sensing elements, and the heat relay,

2.18.6.6

(ET)

Cabin Air Temperature

bypassing the air conditioning discharge sensing

Control. The cabin air temperature control panel

element. When the bridge is balanced, the polarized

(Figure 2-50.2) is mounted on the bulkhead behind the

switch is de-energized, cutting power to the control unit

copilot. It is a variable resistor which varies the

heat relay, which, in turn, de-energizes the fuel control

resistance to unbalance the control unit bridge until the

valve solenoid to stop the heater. When the bridge is

desired cockpit and cabin temperature is achieved. For

unbalanced, current flow across the bridge energizes the

maximum cooling, the knob may be rotated to

polarized switch allowing electrical power to flow to

RECIRCULATE. This closes the outside air damper

the heat relay, which in turn energizes the fuel control

and recirculates the cabin air through the air conditioner

valve solenoid, starting the heater cycle. With the air

system.

conditioning system operating, the control relay

energizes and unbalances the bridge. At this time, the

2.18.6.7

(ET) Sensing Elements. Five sensing

energized control relay cuts current to the heater

elements form part of the control unit bridge circuit.

discharge sensor and permits current flow through the

These elements include: two cabin sensing elements in

CABIN AIR TEMPERATURE rheostat, outside air

the cabin; a heater discharge sensing element on the

temperature sensing element, cabin sensing elements,

plenum chamber; an outside air temperature sensing

and the air conditioning sensing element. When the

element in the heater intake duct; and an air conditioner

bridge is unbalanced, the polarized switch is energized,

discharge sensing element in the air conditioner duct

allowing electrical power to flow to the motor-

which extends through the cabin floor at station 180.

compressor.

Each compartment-sensing element includes two

2-79

ORIGINAL

NAVAIR 01-230HLH-1

HEAT

VENT

EVAP

MODE

AIR

SELECTOR

BLO

MAN

HEATER

OVRD

CABIN AIR TEMPERATURE

COLD

WARM

OFF

AIR

AUTO

RECIRCULATE

COND

01771028

Figure 2-50.2. (ET) Air Conditioning Control and Cabin Air Temperature Controls

2.18.6.9(ET) Air Distribution System. The distribution

2.18.6.10 (ET) Heating System. The heating

system consists of a manually operated damper, electrically

system is similar to that installed on the UH-3H

operated damper, actuator, automatic damper, air

helicopters, except that it can operate as an automatic

conditioning blower, heater blower, diffusers and registers,

thermostatically controlled system.

Refer to the

plenum chamber, transition duct, and the necessary heating

following subparagraphs for specific differences. Refer

and ventilating ducts to conduct the flow of air through the

to NAVAIR

01-230HLH-2-1.1 for complete system

cockpit and cabin area. A manually operated damper

description and theory of operation.

governing air conditioning or ventilating air for the cabin is

mounted on the ductwork, under the one-place seat, on the

2.18.6.10.1

(ET) Heater. The heater is similar to

right side of the forward entrance compartment. The

that installed on the UH-3H helicopters, except that it is

electrically operated damper, controlled by the actuator, is

a 200,000 BTU heater, the exhaust is on the bottom of

in the forward tub compartment and is a normally open

the heater instead of at the center, and the spark plug is

damper. During recirculation of interior air, the damper

at a 90-degree angle to the ground electrode.

remains in the closed position. The automatic damper, in

the forward tub compartment, automatically controls the

2.18.6.10.2

(ET) Plenum Chamber. The plenum

flow of either heated air or air-conditioned air into the cabin

chamber is similar to that of the heater installed on the

area, depending upon which system is in operation. The air-

UH-3H helicopters, except that the

65.6°C

(150°F)

conditioning blower, in the forward tub compartment, draws

thermal switch has been removed and the heater

outside air into the helicopter for the air conditioning system

discharge-sensing element has been installed in its

through an intake vent forward of the forward passenger

place.

door. The blower is also utilized to recirculate interior air,

at which time the electrically operated damper is closed.

2.18.6.10.3

(ET) Thermal Switches. The thermal

The heater blower draws outside air and forces it through

switches are similar to those of the heater installed on

the otherwise inoperative heating system and into the

the UH-3H helicopters, except that the 65.6°C (150°F)

helicopter. The overhead ducts, on both sides of the cabin

switch is not installed.

ceiling, include individual units called gaspers. The gaspers

control the flow of ventilating or air conditioning air into the

2.18.6.11

(ET) Air Conditioning S ystem. The

cabin for individual passenger comfort. Two heating,

air conditioning system is a vapor cycle system that

ventilating, and air conditioning ducts, connected to the

provides a comfortable environment for crew members

manually operated damper, run forward under the cockpit

and passengers.

floor to registers and diffusers on both sides of the cockpit.

A single heating, ventilating, and air conditioning duct,

connected to the plenum chambers by a transition duct, runs

aft along the right-hand side of the cabin. Two circular

adjustable and eight nonadjustable diffusers in the cabin

disperse heated, ventilated, and air conditioned air.

2-80

ORIGINAL

NAVAIR 01-230HLH-1

2.18.6.11.1 (ET) Evaporator Pallet Assembly. The

duct. The demister assembly consists of an aluminum

evaporator pallet assembly is an aluminum honeycomb

housing, a fine aluminum wire mesh demister pad mounted

pallet located below the cabin floor at station

180,

between a screen and two mounting brackets, and a drain

horizontally-mounted on the centerline of the aircraft.

tube for the removal of condensation. The demister

Major components mounted to the pallet are:

removes moisture from the conditioned air. The moisture

condenses and collects on the wires of the demister pad

Compressor/motor Assembly

forming droplets. The droplets are pulled by gravity down

Evaporator Fan

to the drain pan where they are then dumped overboard

Inlet Transition Duct

through the drain tubes.

Heat Exchanger

Demister Assembly

2.18.6.11.1.6

(ET) Low Temperature Switch. The

Temperature Sensors

low temperature switch is mounted on the aft side of the

Hot Gas Bypass Valve

demister housing. The low temperature switch will trip the

low temp fault indicator and disengage the compressor if the

2.18.6.11.1.1(ET) Compressor/motor

Assembly.

evaporator outlet temperature drops below -1.1 - 4.4°C (30-

The compressor/motor, mounted in the forward tub

40°F). The system will automatically resume normal operation

compartment, compresses and circulates the refrigerant in

when the temperature rises to 12.8°C (55°F).

the air conditioner system. The compressor/motor assembly

is a single, hermetically sealed unit, requiring 115/200-volt,

2.18.6.11.1.7

(ET) Hot Gas By-pass Valve. The hot

400 Hz, three-phase ac current for operation. A thermally

gas by-pass valve is located on the evaporator pallet next to

protected six horsepower motor is mounted vertically over a

the heat exchanger. The valve provides regulation of the

vane rotary type pump that compresses and circulates

evaporator outlet air temperature. By adjusting the cockpit

refrigerant throughout the system.

Mounted on the

temperature control, through the temperature controller in

aluminum compressor/motor housing is a connector for the

the electronic box, the hot gas by-pass valve will discharge

electrical power, refrigerant discharge and suction ports, and

hot refrigerant gas from the compressor discharge tube

an oil level sight glass. The No. 1 generator, through the

directly into the heat exchanger. The valve receives input

35-ampere AIR COND COMPR MOTOR circuit breaker on

from the low temperature switch mounted on the demister

the copilot’s circuit breaker panel, provides power for the

housing and the temperature controller. The valve bypasses

motor-compressor.

the expansion valve with hot gas regulating the evaporator

pressure; with an increase in refrigerant pressure the

2.18.6.11.1.2 (ET) Evaporator Fan. The evaporator fan

temperature is also increased.

is a

6-inch diameter, two-stage axial vane type fan with

internal motor windings, and an explosion proof housing. It

2.18.6.11.1.8

(ET) Thermal Electric Expansion

requires

115/200-volt ac, three-phase,

400 Hz power for

Valve. The thermal electric expansion valve meters the

operation. The fan draws return air from the cabin interior

refrigerant flow into the evaporator heat exchanger during

and circulates it through the evaporator heat exchanger. The

the air conditioner operation. The orifice flow of the valve

fan is located between the cabin return duct and the inlet

is controlled by voltage applied to the electrical terminals on

transition duct.

the valve head. The valve incorporates a resistance wire on

the valve head. The valve incorporates a resistance wire-

wound, bimetal motor. The valve is operated by, and

2.18.6.11.1.3

(ET) Inlet Transition Duct. The inlet

responds to, low voltage electrical power. The amount of

transition duct is located between the evaporator and the

electrical power applied to the valve controls the degree of

heat exchanger. The duct directs the flow of air from the

valve opening. At zero voltage, the valve is closed. As

fan outlet to the heat exchanger inlet. It is constructed of

voltage is applied, heat deflects the bimetal motor causing

three layers of epoxy impregnated fiberglass laminate.

the valve needle to follow.

2.18.6.11.1.4

(ET) Evaporator Heat Exchanger.

2.18.6.11.1.9

(ET) Liquid Sensing Thermister.

The evaporator heat exchanger is located between the inlet

The liquid sensing thermister is located in the evaporator

transition duct and the demister assembly. It is of a plate fin

suction line and regulates the voltage to the thermal electric

design and constructed of aluminum. The heat exchanger

expansion valve. The thermister reacts to the refrigerant,

transfers heat from the cabin air to the refrigerant.

increasing or decreasing the voltage to the valve, depending

on the state of the refrigerant exiting the evaporator. Liquid

2.18.6.11.1.5

(ET) Demister Assembly.

The

closes the valve and gas opens the valve.

demister assembly is mounted on the outlet of the heat

exchanger and provides a mounting surface for the outlet

2 - 81

ORIGINAL

NAVAIR 01-230HLH-1

2.18.6.11.2

(ET) Condenser Assembly.

The

2.18.6.11.3.1

(ET) Filter Dehydrator. The filter

condenser assembly is mounted horizontally in the left

dehydrator is mounted on the auxiliary service pallet in the

sponson. Refrigerant, after being heated and brought to a

high pressure (liquid) refrigerant line between the condenser

high pressure by the compressor, passes through the

and the expansion valve. Contaminants and moisture are

condenser where cooling air condenses the gas. Major

removed from the refrigerant through the filter dehydrator’s

components of the condenser assembly are as follows:

molded porous core.

2.18.6.11.3.2

(ET) Refrigerant Liquid Line

Heat Exchanger

Indicator. The liquid line indicator, or sight glass, is

Transition Duct

located in the liquid line between the condenser heat

Condenser Fan

exchanger outlet and the filter/dehydrator. Its primary

Check Valve

purpose is to allow visible evidence of liquid

clear) or

High Pressure Relief Valve

vapor (bubbles) in the liquid line to determine the state of

the system’s refrigerant charge. The sight glass is not used

2.18.6.11.2.1

(ET) Heat Exchanger.

The heat

to determine the adequacy of the refrigerant charge. The

exchanger is located in the aft portion of the left sponson. It

liquid line indicator also contains a moisture indicator. The

is of a plate fin design and constructed of aluminum to

indicator turns yellow if moisture is in the system refrigerant

minimize weight. The heat exchanger transfers heat from

and green when the system is dry.

the refrigerant to the ambient air.

2.18.6.11.3.3.

(ET) High Pressure Switch. The

2.18.6.11.2.2

(ET) Transition Duct. The transition

high-pressure switch will disengage the compressor if

duct connects the heat exchanger and fan. It is attached to

system refrigerant pressure reaches

350 ± 20 psi. The

flanges on the components and is supported by brackets.

compressor will reengage as refrigerant pressure decreases

to 150 psig. The high-pressure switch is mounted on the

2.18.6.11.2.3

(ET) Condenser Fan. The condenser

refrigerant servicing manifold on the auxiliary service

fan draws ambient air through the heat exchanger and

pallet.

exhausts it overboard. The 12-inch axial vane fan motor

requires

115/200 volts,

400 Hz, three-phase electrical

2.18.6.11.3.4

(ET) Low Pressure Switch. The low-

current for operation. The fan is explosion proof and

pressure switch, in the event of complete refrigerant loss,

thermally protected.

will disengage the compressor when the system pressure

drops below 50 ± 3 psig. When the system pressure rises

2.18.6.11.2.4

(ET) Check Valve. The check valve is

above 65 psig, the compressor will reactivate. The low-

mounted on the refrigerant inlet port of the heat exchanger.

pressure switch is mounted on the refrigerant-servicing

It is a poppet-type valve that prevents a back flow of

manifold on the auxiliary service pallet.

refrigerant when the system is shut down.

2.18.6.11.4

(ET)

Electrical

Control

Box

2.18.6.11.2.5

(ET) High Pressure Relief Valve.

Assembly. The electrical control box contains all the

The high pressure relief valve is located in the high pressure

major electrical components necessary for air conditioning

liquid line and is designed to vent refrigerant from the

operation. The control box is mounted in the forward tub

system in the event that the high pressure switch fails and

compartment and includes the following components:

the system refrigerant pressure exceeds 475 psig. The valve

vents refrigerant overboard. As pressure decreases below

Control Relays

475 psig, the valve will automatically reseal and normal air

Time Delays

conditioning operation will continue.

Temperature Controller

Fault Indicator Panel

2.18.6.11.3

(ET) Auxiliary Servicing Pallet. The

auxiliary servicing pallet is an aluminum honeycomb pallet,

2.18.6.11.4.1

(ET) Control Relays. The three phase

located in the forward tub compartment under the cabin

electrical power is controlled by power contactor type

floor. Major components of the auxiliary servicing pallet

relays.

One relay is used for each of the major air

are:

conditioning system components.

Filter Dehydrator

2.18.6.11.4.2

(ET) Time Delays. Two time delays are

Refrigerant Liquid Line Indicator (Sight Glass)

used in the control of the air conditioning system to prevent

High Pressure Switch

all of the electrical loads from starting simultaneously,

Low Pressure Switch

thereby reducing the electrical surge. A five-second-time

Refrigerant Servicing Ports

2 - 82

ORIGINAL

NAVAIR 01-230HLH-1

delay controls the power contactor for the condenser fan. A

2.19 ANTI-ICING SYSTEM

ten-second-time delay relay controls the power contactor for

the compressor.

The anti-icing system

(Figure

2-51) provides ice

protection for the engine air inlets, the engine starter

2.18.6.11.4.3

(ET) Temperature Controller. The

fairings, and the windshield. The engine air inlets are anti-

temperature control unit contains solid-state circuits

iced by electrically heating the intake duct, thus preventing

necessary for thermostatic temperature control of the air

ice buildup that would cause shedding of ice into the

conditioning system. The controller requires two inputs for

compressor. Compressor bleed air provides anti-icing of the

operation; one input from the rotary temperature selector set

engine starter fairing and mounting struts. Electrical heating

at the desired cabin temperature and one input from the

of the windshield provides deicing of both the pilot and

supply air temperature sensor. In the air conditioning mode

copilot windshield.

the temperature controller cycles the hot gas bypass valve

on and off to control the outlet temperature.

2.19.1 Engine Air Inlet Anti-Icing System. The engine

air inlets are anti-iced by thermal electric resistance

2.18.6.11.4.4

(ET) Fault Indicator Panel. The fault

elements embedded within the epoxy glass intake ducts.

indicator is mounted in the electrical control box cover to

Electrical current is applied to the resistance elements to

indicate extremes in either pressure or temperature during

heat the duct skin higher than the temperature at which ice

air conditioning operation. The panel consists of three one-

will form for OATs above -18^o C. The engine air inlet anti-

amp circuit breakers. The fault indicators are labeled High

icing system should be turned on when operating in

Pressure, Low Pressure, and Low Temperature. The system

conditions of 10^oC OAT and below where visible moisture

will automatically resume operation after a fault condition

(rain, fog, clouds, etc.), ice, or snow accumulation is noted

subsides. The affected circuit breaker should be reset prior

on the helicopter. Significant locations to be observed for

to the next flight.

ice or snow accumulation include the windshield wipers, the

pitot tubes, and the stub wings. The engine air inlet anti-ice

2.18.6.11.5

(ET) Remote Service Manifold. The

should remain off when flying in dry snow and no

remote service manifold assembly is installed at station 155

accumulation is noted on the helicopter. The engine anti-ice

on the left side of the aircraft. Removal of an external

switches (Figure 2-52) on the overhead switch panel (Figure

access panel permits external servicing of the air

2-9) turn the system on. The inlet temperature is limited to a

conditioning system.

maximum of 93^o C by an overheat sensor, embedded in the

intake duct, and recycled on at a minimum of 81^oC by the

2.18.6.11.6

(ET) Ventilating System.

engine controller. The resistance elements of the Nos. 1 and

ventilating system is incorporated in the heating and air

2 engines are operated by ac power and 28 vdc provides

conditioning system. It provides circulation of outside air

control.

throughout the inside of the helicopter. Positioning the

VENT AIR switch, on the air conditioning control panel, to

ON actuates the system. This energizes the air conditioning

and heater blower relays that start the blowers.

Figure 2-51. Anti-Icing System

2 - 83

ORIGINAL

NAVAIR 01-230HLH-1

The thermal switch in the air inlet duct closes to connect

closed

(energized) position by

28-vdc power. This

dc power to the CAUTION PANEL, lighting the #1 INLET

percent loss will occur only if the engine is operating at

ANTI-ICE and/or #2 INLET ANTI-ICE when the air inlet

topping power. With complete dc power failure, the

temperature drops below 38^oC with the ENGINE ANTI-

solenoid valve will open but the advisory panel light will

ICE switch ON. If OAT is -18 ^oC or above, and the engine

not go on.

anti-icing system is on, the CAUTION PANEL lights

coming on will indicate anti-ice system failure. If OAT is

2.19.3 Engine Anti-Ice Switches. Two engine anti-

-18^o C or below and the engine anti-ice system is on, the

ice switches (Figure

2-52) marked ANTI-ICE ENGINE 1

CAUTION PANEL lights will go on, indicating no anti-ice

and 2 with marked positions OFF and ON are on the

protection, as electrical heating is not enough to maintain

overhead switch panel

(Figure

2-9). ON energizes the

38^oC under these ambient conditions. However, as previously

engine air inlet anti-icing system and activates the engine

mentioned, the engine air inlet anti-icing system should not

anti-icing system. OFF turns the anti-icing systems off. If

be turned on unless there is a visible moisture or ice or

the ambient temperature is

38^o C or below, placing the

snow accumulation. (For a more detailed discussion refer

engine anti-ice switches ON turns on the engine anti-ice

to Extreme Weather Operations, Chapter 14.)

and engine air inlet anti-ice systems and lights the #1

INLET ANTI-ICE and

#2 INLET ANTI-ICE caution

2.19.2 Engine Anti-Icing System. The engine starter

lights and the advisory panel lights marked #1 ENG ANTI-

fairing (Figure 2-51), the inlet guide vanes, and the top,

ICE and #2 ENG ANTI-ICE. When the engine air inlet

right, and left struts of the front frames of each engine are

anti-icing system has increased the duct air temperature to

anti-iced by diverting engine tenth-stage compressor air to

38^oC, the caution panel lights will go off.

heat them. Actuating the engine anti-ice switch (Figure 2-

52) on the overhead switch panel (Figure 2-9) deenergizes

2.19.4 Windshield Anti-Ice System. The pilot and

an engine-mounted solenoid valve to the open position,

copilot windshields are anti-iced by an electric current that

allowing hot compressor air to flow through the engine

passes through a transparent resistant film on the inner

front frame to the inside of the starter fairing and the inlet

surface of the outer pane of the windshield (Figure 2-51).

guide vanes. When the engine solenoid is deenergized,

The windshield anti-icing system consists of the treated

lights are illuminated on the advisory panel, displaying to

windshields, a windshield anti-ice controller, transformers,

the pilot the markings #1 ENG ANTI-ICE ON and/or # 2

and a switch on the overhead switch panel. The windshield

ENGINE ANTI-ICE ON. The engine anti-icing system is

anti-icing also serves to defog the windshield.

turned on simultaneously with the engine air inlet anti-

icing system. If the engine anti-ice advisory light on the

2.19.4.1 Windshield Anti-Icing Switch. This switch

advisory panel lights during flight when the engine anti-ice

(Figure

2-53) marked WINDSHIELD with marked

switches are OFF and the anti-icing system is not in

positions LOW, OFF, and NORMAL is on the overhead

operation, it indicates that the engine anti-ice solenoid

switch panel

(Figure

2-9). When placed to LOW, the

valve has opened

(deenergized) because of electrical

windshield controller is energized, current heats the

failure and that a loss of about

6-percent engine

windshield. The low temperature setting is used for

horsepower will occur. The solenoid valves are held in the

defogging and light icing conditions.

Figure 2-52. Engine Anti-Ice Switches

Figure 2-53. Ice Protection Switches

2 - 84

ORIGINAL

UH-3H and UH-3H EXECUTIVE TRANSPORT. FLIGHT MANUAL (2000) - page 2