C-130T. FLIGHT MANUAL (2006)

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CHAPTER 2

Systems and Equipment

2.1

ENGINES

The aircraft is powered by four Allison T56-A-16 turboprop engines (see Figure 2-1). The basic engine consists of

two major assemblies, a power section, and a reduction gear assembly, which are attached to each other by an

extension shaft assembly and two supporting struts. The engine is provided with fuel, oil, starting, ignition, and

control systems. The engine operates at a constant speed; therefore, engine power is related to TIT, which varies

according to the rate of fuel flow. An increase in fuel flow causes an increase in TIT and a corresponding increase

in energy available at the turbine. The turbine then absorbs more energy and transmits it to the propeller in the form

oftorque.In orderto absorbtheincreasedtorque, thepropellerincreasesbladeangleto maintainconstant enginerpm.

A decrease in torque results in a decrease in propeller blade angle to maintain engine speed. Thrust is obtained from

the propeller, and a small amount of additional thrust (approximately 10 percent at takeoff) is created by the tailpipe

exhaust.

2.1.1 Power Section

The power section of the engine is composed of a single-entry, 14-stage, axial-flow compressor; a set of six

combustion chambers of the through-flow type; and a 4-stage turbine. Mounted on the power section are an

accessories drive assembly and components of the engine fuel, ignition, and control systems. Acceleration bleed

valves are installed at the 5th and 10th compressor stages. A manifold is installed at the diffuser to bleed air from

the compressor for aircraft pneumatic systems. Anti-icing systems are provided to prevent accumulation of ice in the

engine inlet air duct and the oil cooler scoop. Inlet air enters the compressor through a scoop and duct below the

compressor and is progressively compressed through the 14 stages of compression. The compressed air (at

approximately 125 psi and 600 _F) flows through a diffuser into the combustion section. Fuel is introduced into the

combustion chambers and burned to increase the temperature and, thereby, the energy of the gases. The gases pass

through the turbine causing it to rotate and drive the compressor, the propeller, and accessories. The gases, after

expanding through the turbine, flow out through a tailpipe.

2.1.2 Extension Shaft Assembly

The extension shaft assembly consists of two concentric shafts and torquemeter components. The inner shaft

transmits power from the power section to the reduction gear. The outer shaft serves as a reference so that the torsional

deflection of the loaded inner shaft can be detected by the magnetic pickups of the torque indicating system.

Torquemeter amplifiers are provided with adjustment screws for calibration purposes.

2.1.3 Reduction Gear Assembly

The reduction gear assembly contains a reduction gear train, a propeller brake, a NTS system, and a safety coupling.

Mounted on the accessory drive pads are the engine starter, an ac generator, a hydraulic pump, an oil pump, and a

tachometer generator. The reduction gear has an independent dry-sump oil system. The reduction gear train is in two

stages, providing an overall reduction of 13.54 to 1 between engine speed (13,820 rpm) and propeller shaft speed

(1,021 rpm). The propeller brake, NTS system, and safety coupling are described in the following paragraphs.

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2.1.3.1 Propeller Brake

The cone-type propeller brake acts on the first stage of reduction gearing. During engine operation, it is held

disengaged by gearbox oil pressure when rpm exceeds 23 percent and is engaged below this speed. As engine speed

is reduced and oil pressure drops, the braking surfaces are brought into contact by spring force to help slow the

propeller to a stop. Helical splines between the starter shaft and the starter gear on the outer brake member cause the

brake to disengage when starting torque is applied during starting. The brake also engages to stop reverse rotation

of the propeller.

2.1.3.2 Negative Torque Protective Devices

Very high windmilling drag upon sudden loss of turbine power is a problem in propeller-jet aircraft operation. High

windmilling drag results when the compressor absorbs a great amount of power. If a power failure is experienced

in flight at high speed, the engine starts to slow down; the propeller, sensing rpm, reduces blade angle and drives the

compressor, trying to bring it up to speed. Two safety devices are provided to prevent a sudden increase in

windmilling drag that could result in high structural loads on the tail of the aircraft and loss of directional control.

These devices are the NTS system, which increases blade angle of the propeller, and the safety coupling, which can

decouple the propeller and reduction gear from the power section of the engine upon failure of the power section.

2.1.3.3 Negative Torque Signal System

The NTS system provides a mechanical signal to limit negative torque. Negative torque is encountered when the

propeller attempts to drive the engine. If not relieved, this condition creates a great amount of drag, causing the aircraft

to yaw. The NTS system consists of an actuating mechanism housed partly within the reduction gear assembly and

partly in the propeller control assembly. It operates when negative torque applied to the reduction gear exceeds

approximately 1,260 ±600 inch-pounds. A ring gear is then moved forward against springs as a result of torque

reaction generated through helical splines. In moving forward, the ring gear actuates a plunger extending through

thenoseofthegearbox.Theplungerpushesagainstacaminthecontrolassemblytoactuatecontrollinkageconnected

tothepropellerfeathervalve.WhenaNTSistransmittedtothepropeller,thepropellerincreasesbladeangletorelieve

thecondition, except when thethrottles arebelow theFLIGHT IDLEposition. WhenthethrottlesarebelowFLIGHT

IDLE, a cam moves the actuator away from the NTS plunger and renders the system inoperative. This is necessary

to prevent a propeller from receiving a possible NTS at high landing speeds when the throttles are moved toward

reverse. If the negative torque is sufficiently reduced, the signal mechanism returns to normal by springs acting on

the ring gear.

Note

Properly adjusted, the NTS will not commit the propeller to feather. An

improperly adjusted or malfunctioning NTS system receiving continual

NTSs will cause a propeller to move to a high blade angle. The resultant

load may cause an engine stall and flameout.

2.1.3.4 Safety Coupling

The safety coupling is provided to decouple the power section from the reduction gear if negative torque applied to

the reduction gear exceeds approximately 6,000 inch-pounds, a value much higher than that required to operate the

NTS system. Because of its higher setting, the safety coupling backs up the NTS system to reduce drag until the

propeller can be feathered. The safety coupling connects the engine extension shaft to the pinion of the first stage

ofreduction gears. It consists ofthree members: an outermember is attached to the extension shaft; an innermember

is attached to the pinion; and an intermediate member is engaged to the outer member by straight teeth and to the

inner member by helical teeth. Reaction of the helical teeth tends to force the intermediate member aft out of

engagement when negative torque is applied, and the members disengage if approximately 6,000 inch-pounds of

negative torque is reached. While disengaged, the two members are forced together by springs so that the teeth will

ratchet. The teeth can thus be damaged; therefore, the engine should not be continued in operation after a decoupling.

Before restarting the engine, the coupling must be replaced.

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2.1.4 Engine Fuel and Control System

In flight, the engine operates at a constant speed that is maintained by the governing action of the propeller. Power

changes are made by changing fuel flow and propeller blade angle rather than engine speed. An increase in fuel flow

causes an increase in TIT and a corresponding increase in energy available at the turbine. The turbine absorbs more

energy and transmits it to the propeller in the form of torque. The propeller, in order to maintain governing speed,

increases blade angle to absorb the increased torque. TIT is a very important factor in the control of the engine. It

is directly related to fuel flow and consequently to power produced. It is also limited because of the strength and

durability of the combustion and turbine section materials. The fuel control system schedules fuel flow (see Figure

2-2) to produce specific TITs and to limit those temperatures so that the temperature tolerances of combustion and

turbine section materials are not exceeded. Changes in power settings are effected by the throttle, which is connected

to the fuel control and the propeller through a mechanical coordinator. During ground operation, changes in throttle

position mechanically affect both the fuel flow and the propeller blade angle. In flight, changes in throttle position

mechanically affect fuel flow, and the propeller governor regulates blade angle, maintaining constant engine speed.

The hydromechanical fuel control, which is part of the basic fuel system, senses engine inlet air temperature and

pressure, rpm, and throttle position, and varies fuel flow accordingly. The electronic TD control system senses TIT

and throttle position and makes any necessary changes in the fuel flow from the fuel control before it reaches the fuel

nozzles. The TD system compensates for minor variables not sensed by the hydromechanical fuel control and for

mechanical tolerances within the fuel control itself. By means of switches, the TD system can be turned off (NULL

position) and the engine will operate on the basic hydromechanical system alone. With the TD system in AUTO,

temperature protection is provided through the entire throttle range and automatic temperature scheduling is provided

when the throttle is in the range of 65_ to 90_. When the TD system is in NULL, the functions of temperature limiting

and temperature scheduling must be accomplished manually by adjustment of the throttle.

2.1.4.1 Basic Hydromechanical Fuel System

Thebasichydromechanicalfuelsystemconsistsofathrottle,acoordinator,alow-pressurefuel filter,a high-pressure

fuel filter, a dual-element fuel pump, a hydromechanical fuel control, and six fuel nozzles.

2.1.4.2 Throttle, Coordinator, and Propeller Control Linkage

The coordinator is a mechanical discriminating device that coordinates the throttle, the propeller, the fuel control,

and the electronic TD system. Movements of the throttle are transmitted to the coordinator and, in turn, to the fuel

control and the propeller by a series of levers and rods. A potentiometer in the coordinator provides signals to the

TD system. Propeller blade angle is scheduled by throttle position from MAXIMUM REVERSE to FLIGHT IDLE.

For throttle settings between FLIGHT IDLE and TAKE-OFF, the propeller is governing. Throttle movement in this

range serves primarily to change fuel flow and also to change propeller hydraulic pitch stop (beta followup) settings.

2.1.4.3 Fuel Control

Fuel flows from the fuel pump to the hydromechanical fuel control. The control is sensitive to throttle position, air

temperature and pressure at the engine inlet, and engine speed. The engine speed function of the fuel control maintains

engine speed in the taxi range and limits engine speed in the flight range if the propeller governor fails. Governor

action is controlled by flyweights that respond to engine rpm.

Thecontrol will start to reducefuel to theengineat approximately 103.5 percent at a rateof 900 pounds per1-percent

rpm. A speed servo system provides for speed acceleration during engine start and underspeed control during normal

engine operation. Fuel metered by the control is equal to engine requirements plus an additional 20 percent, which

is fortheuseoftheTD valve. With theTD system in NULL, theexcess fuel provided by the fuel control is constantly

bypassed by the TD valve back to the fuel pump, and fuel metering is accomplished by the fuel control alone. The

required fuel flow passes on through the TD valve to the fuel nozzles and into the combustion liners, where it is

burned.

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2.1.4.4 Fuel Pump and Filter Assembly

Thepumpandfilterassemblycontainsacentrifugal boostpump, aprimary andasecondarygearpump,high-pressure

filter, pressure switch, pump paralleling valve, and bypass and check valves. After engine start, the primary and

secondary gear pumps operate in series; that is, the secondary pump receives fuel from the centrifugal pump and

directs its output to the primary pump, which supplies fuel to the engine. If either pump fails, the operating pump

will bypass the failed pump and supply sufficient fuel to sustain normal full-power engine operation. During engine

starting, the solenoid-operated pump paralleling valve is energized closed by the speed-sensitivecontrol through the

ignition relay to block flow from the secondary to the primary pump. The centrifugal boost pump then feeds both

gear pumps, and the combined output of the pumps operating in parallel is fed to the engine to provide increased fuel

flow capacity during the start cycle. The pressure switch is closed to turn on the secondary pump pressure light while

the pump paralleling valve is closed, or when the primary pump has failed.

2.1.4.5 Low-Pressure Filter

A low-pressure, paper-element filter assembly is installed between the centrifugal boost pump and the two gear

pumps to remove contamination from the incoming fuel. This filter assembly is mounted externally from the fuel

pump and filter assembly.

2.1.4.6 Acceleration Bleed-Air Valves

The bleed-air valves on the 5th and 10th stages of the compressor are provided for compressor unloading during

starting and while the engine is operating in the low-speed, ground-idle range. These bleed valves remain open only

when engine speed is below 94-percent rpm. The 5th- and 10th-stage bleed-air valves are automatic in operation and

are actuated by 14th-stage compressor air pressure through an engine-driven, speed-sensitive valve assembly.

2.1.4.7 Starting Fuel Enrichment System

The enrichment system consists of a bypass line in which is mounted a solenoid valve, a pressure switch, and the

engine fuel-enrichment selector switch. The valve is opened by the speed-sensitive control through the ignition relay

when engine speed reaches 16-percent rpm during starting when the engine fuel enrichment switch is in NORM.

While open, it allows pump discharge fuel to flow around the metering section of the fuel control to add to the metered

flow from the fuel control. After fuel pressure in the manifold reaches approximately 50 psi (gauge), the manifold

pressure switch opens to deenergize the valve, which then closes.

2.1.4.8 Fuel Nozzles

Six duplex fuel nozzles are used. The small slots of the nozzles open when manifold pressure is low during engine

starting. Both small and large slots open when manifold pressure increases to 65 psi as the engine accelerates to

normal speed.

2.1.4.9 Burner Drain Valves

Two burnerdrain valvesareprovidedto drainfuel fromthecombustionsection oftheenginewhen itis stopped.They

are spring-loaded and operate automatically to close when internal combustion pressure exceeds external pressure.

2.1.4.10 Drip Valves

Thepurposeofthedripvalveistodrainthefuelmanifoldtominimizetheamountoffueldrippingintothecombustion

chambers and causing fire at shutdown. The valve is closed by a solenoid that is controlled by the speed-sensitive

control through the ignition relay when engine speed reaches approximately 16-percent rpm during starting. It is held

against a spring by pressure in the fuel manifold after the engine is started. When fuel pressure in the manifold drops

to approximately 9 psi during engine shutdown, the valve opens and drains the manifold.

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2.1.5 Starting System

An airturbinestarterunit drives theengineforground starts.This starterunit consistsofanair-driven turbinesection,

a clutch, and a reduction gear section that is splined to the reduction gear assembly of the engine. Air for driving the

starter can be supplied by the auxiliary power unit, by an operating engine, or by an external air source. The air is

routed through the bleed-air system and the engine bleed-air regulators.

When the respective bleed-air regulator is opened, air is supplied to the starter regulator valve. When the ENGINE

GROUND START switch is placed to START, the starter regulator valve opens (when its solenoid is energized) and

allows airflow into the starter turbine section. Releasing the ENGINE GROUND START switch to OFF will

deenergize the regulator valve. Within approximately 15 seconds the START VALVE OPEN light will extinguish,

indicating that the starter valve is closed. Each engine starting circuit is electrically interlocked with the

corresponding engine oil fire shutoff valve control circuit. This renders the starting circuit inoperative unless the fire

emergency control handle is pushed in and the FIRE SHUTOFF VALVE-OIL circuit breaker is engaged.

2.1.6 Ignition System

The ignition system is a high-voltage, condenser-discharge type, consisting of an exciter, two igniters, and control

components. The system is controlled by the speed-sensitive control through the ignition relay, which turns it on at

16-percent engine rpm and off at 65-percent engine rpm during starting.

2.1.6.1 Electronic Temperature Datum Control System

The electronic TD control, together with the coordinator potentiometer, temperature adjustment network, a TIT

measurement system, and the temperature datum valve make up the electronic temperature datum system. The system

compensates for variations in fuel heat value and density, engines, and control system characteristics. The TD control

is furnished actual TIT signals from a set of thermocouples, and is furnished desired TIT signals by the throttle

through thecoordinator potentiometerand thetemperature adjustment network. Thecontrol compares the actual and

the desired TIT signals. In the temperature controlling range (65_ to 90_ throttle movement), if there is a difference,

the TD control signals the temperature datum valve to increase or decrease fuel flow to bring the temperature back

on schedule. In the temperature limiting range (0_ to 65_ throttle movement), the TD control acts only when the

limiting temperatureis exceeded, at which time, it signals theTD valveto decreasefuel flow.TheTDvalveislocated

between the fuel control and the fuel nozzles. It is a motor-operated bypass valve that responds to signals received

from the TD control. In throttle positions between 0_ and 65_, the valve remains in a 20-percent bypass or null

position and theengineoperates on thefuel flow scheduled by thefuel control. The valveremains in the null position

unless it is signaled by the TD control to limit TIT. The valve then reduces the fuel flow (up to 50 percent during

starting, 20 percent above 94-percent rpm) to the nozzles by returning the excess to the fuel pump. When the TIT

lowers to thedesired level, theTD control signals thevalveto return to thenullposition. Inthrottlepositionsbetween

0_ and 65_, the control system is in the temperature limiting range. In throttle positions between 65_ and 90_, the

TD valve acts to control TIT to a preselected schedule corresponding to throttle position; this is the temperature

controlling range. In this range the valve may be signaled by the TD control to allow more (higher temperature

desired) or allow less (lower temperature desired) of the fuel to flow to the fuel nozzles. Any specific fuel flow trim

correction applied in the 65_ to 90_ throttle range can be locked into the TD valve while above 65_ and will be

maintained in the 0_ to 65_ range by the use of the TEMP DATUM CONTROL VALVE switch located at the flight

station.Also,theTDsystemcanbereturnedtonullatanytimebytheuseoftheTEMPDATUM CONTROLVALVE

switch. When the switch is in NULL, automatic temperature-limiting circuits are inoperative, the TD valve remains

in the null (20-percent bypass) position, and all fuel metering is then accomplished by the fuel control. Temperature

limiting then must be accomplished by throttle adjustment.

2.1.7 Engine Controls and Control Systems

Engine control in the flight range of operation is based on regulation of engine speed by propeller constant-speed

governing and control of torque through regulation of fuel flow. Note that the throttle acts only as a power control.

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It exercises no direct control over the propeller, which is controlled entirely by the propeller regulator to regulate

engine speed and to limit the low blade angle. The fuel control regulates the rate of increase and decrease of fuel

metering for acceleration and deceleration. The TD control system functions at all throttle positions (MAXIMUM

REVERSE through TAKE-OFF) either as a temperature-limiting or a temperature-controlling system. When

performing temperature limiting, the system operates to prevent overtemperatures by “taking” a portion of fuel flow

when maximum allowable TIT is exceeded. When performing temperature controlling (past 65_ of throttle travel,

fuel correction lights out), it adjusts fuel flow to obtain the desired temperature. This is accomplished by using a

voltage signal from a potentiometer, positioned by the throttle, as a reference for desired TIT. The system compares

this reference with actual TIT and adjusts fuel flow to obtain the desired temperature. The TD control system may

be operated to provide a fixed correction of fuel flow. By positioning the TEMP DATUM CONTROL VALVE

switches in LOCKED while the throttles are in the temperature-controlling range, all of the TD control valves are

locked in theposition in which theyhavebeenset bythetemperature-controllingsystems. WhentheTEMPDATUM

CONTROL VALVE switches are positioned in LOCKED, the electronic fuel correction warning lights should

remain out through all throttle movements. Should an overtemperature condition occur, the light, or lights, will come

on indicating that a“take”signal has been initiated, thattheTDvalvebrakehas beenreleased, andthat fuelcorrection

has been lost. The light will stay on until the TD system is reset. In throttle positions below FLIGHT IDLE, the throttle

selects propeller blade angle as well as fuel flow. The TD control system still functions to limit TIT.

An engine is stopped by closing a fuel shutoff valve on the fuel control. The valve is closed electrically when the

condition lever is at GROUND STOP if the control circuit is completed through a landing gear touchdown switch,

when the condition lever is pulled to FEATHER, or when the fire emergency handle is pulled. The valve is also closed

mechanically when the condition lever is pulled to FEATHER. Propeller feathering can be accomplished by pulling

the condition lever to FEATHER, or by pulling the fire emergency handle.

For starting an engine, a speed-sensitive valve and a speed-sensitive control, which contains three speed-sensitive

microswitches,provideautomaticcontrolofallfunctions involvedin thestarting cycle.Thespeed-sensitiveswitches

turn on the fuel, control ignition, parallel fuel pump elements, initiate starting fuel enrichment, close a manifold drip

valve, and switch the electronic datum control system from starting limiting to normal limiting. The speed-sensitive

valve controls the compressor bleed valves. These operations are timed according to engine speed to ensure that the

proper sequence of starting operations is followed. For airstarts, the engine condition lever operates switches that start

the propeller feather pump motor to provide pressure to decrease the propeller blade angle and to apply power to the

speed-sensitive control, energizing fuel and ignition control circuits.

2.1.7.1 Throttles

The throttles (see Figures 2-3 and 2-4) are quadrant mounted on the flight control pedestal. Throttle movements are

transmitted through mechanical linkage to an engine-mounted coordinator. The coordinator transmits the movements

through mechanical linkage to the propeller and to the engine fuel control, and it also actuates switches and a

potentiometer that affect electronic TD control system operation. Each throttle has two distinct ranges ofmovement,

taxi and flight, that are separated by a stop (see Figure 2-3). Both ranges are used for ground operation, but the taxi

range must not be used in flight. In the taxi range, the throttle position selects a propeller blade angle and a

corresponding rate of fuel flow. In the flight (governing) range, throttle position selects a rate of fuel flow and the

propeller governor controls propeller blade angle. The throttles have four placarded positions as follows:

1. MAXIMUM REVERSE (0_ travel) gives maximum reverse thrust with engine power approximately 40

percent of takeoff power.

2. GROUND IDLE (approximately 18_ travel) is a detent position. This is the ground-starting position at which

blade angle is set for minimum thrust.

Note

Throttles must not be moved out of GROUND IDLE detent during ground

starting because the resultant increase in propeller blade angle might

overload the starter, reducing the rate of engine acceleration.

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3. FLIGHT IDLE (34_ travel) is the transition point between the taxi and flight (governing) ranges. A step in

the quadrant limits aft travel of the throttle at this position until the throttle is lifted.

4. TAKE-OFF (90_ travel) is the maximum power position.

The throttle quadrant is also divided into two unmarked ranges with respect to control of the electronic TD

control system. The crossover point is at 65_ throttle travel, at which point the switches in the coordinator are

actuated. Below this point, the electronic TD control system is limiting TIT. Above this point, it is controlling

TIT if the TD valve switches are in the AUTO position.

2.1.7.2 LOW SPEED GROUND IDLE Control Buttons

Four LOW SPEED GROUND IDLE control buttons (see Figure 2-5) on the control pedestal may be pushed in to

reduce engine rpm to approximately 72 percent at any time the throttles are in the range between 9_ and 30_. The

low speed ground idle buttons send a signal to the fuel topping governor on the fuel control, to reduce fuel flow to

the engine. This allows the fuel nozzles to cool down and extends engine life. In addition, this allows the aircraft to

taxi at a slower speed without having to use excess braking. Moving the throttles out of this range will automatically

disengage the LOW SPEED GROUND IDLE buttons. Power is supplied from the essential dc bus through the LOW

SPEED GND IDLE circuit breakers on the copilot side circuit breaker panel.

Figure 2-5. Low-Speed Ground Idle Buttons

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CAUTION

D With engines in low-speed ground idle, movement of throttles beyond the

limits of 9_ to 30_ (coordinator angle) at ambient temperatures above 80

_F may cause engine stall and overtemperature. Therefore, the recom-

mended procedure for coming out of low-speed ground idle is to disengage

the buttons manually with the throttles in GROUND IDLE detent.

D If all four engines are operated at low-speed ground idle, the APU generator

must be on since the engine-driven ac generators will not supply ac power.

If the APU generator fails and the BUS TIE switch is in the TIED position,

the LOW SPEED GROUND IDLE buttons must be disengaged manually

in orderto restoreacpowerand to prevent adrain on thebattery. IftheBUS

TIE switch is not in the TIED position, the LOW SPEED GROUND IDLE

buttons will disengage automatically.

2.1.7.3 Throttle Friction Knob

A friction knob (see Figure 2-4) on the throttle quadrant adjusts the amount of friction applied to the throttles to

prevent creeping or accidental movement.

2.1.7.4 ENGINE CONDITION Levers

Four pedestal-mounted condition levers (see Figure 2-4) are primarily controls for engine starting and stopping and

propeller feathering and unfeathering. They actuate both mechanical linkages and switches that provide electrical

control. Each lever has four placarded positions as follows:

1. RUN is a detent position. At this position, the lever closes a switch that places engine fuel and ignition systems

undercontrolofthespeed-sensitivecontrol.ForenginesNo.2andNo.3,theice-detectionsystem isenergized.

2. AIRSTARTisapositionattainedbyholdingtheleverforward againstspring tension.In thisposition, thelever

closes the same switch closed by placing the lever at RUN and, in addition, closes a switch that causes the

propeller feathering pump to operate.

3. GROUND STOP is a detent position. In this position, the lever actuates a switch that causes the electrical fuel

shutoff valve on the engine fuel control to close only if the landing gear touchdown switches are closed. The

switch also closes the nacelle preheat control circuit making this system operable, if installed.

4. FEATHER is a detent position. When the lever is pulled toward this position, mechanical linkages transmit

themotiontotheengine-mountedcoordinatorandfromthecoordinatortothepropellerandtotheshutoffvalve

on the engine fuel control. Switches are also actuated by the lever as it is pulled aft. The results of moving the

lever to FEATHER are the following:

a. The propeller receives a feather signal and mechanically and electrically energizes the feather solenoid

valve.

b. The fuel shutoff valve on the engine fuel control is closed both mechanically and electrically.

c. The propeller feathering pump is turned on.

d. The nacelle preheat system remains operable only when the aircraft is on the ground (if installed).

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CAUTION

When pulling a condition lever to FEATHER, pull it all the way to the

detent to ensure that the propeller is fully feathered when the engine fuel

is shut off. If the lever is left at midposition and the NTS is inoperative, an

engine decoupling is possible.

2.1.7.5 TEMP DATUM CONTROL VALVE Switches

Four temperature datum control valve switches (see Figure 2-6) are mounted on a control panel on the flight control

pedestal. Each switch has AUTO, LOCKED, and NULL positions. The switch positions are used as follows.

The AUTO position permits normal operation of the electronic TD control system by applying single-phase, ac power

to the amplifier through a FUEL & TEMPERATURE CONTROL circuit breaker on the pilot lower circuit breaker

panel.

The LOCKED position may be set when the throttles are in temperature-controlling range, to provide a

fixed-percentage correction on the metered fuel flow throughout the engine operating range and will permit the fuel

control to compensate for changes in ambient temperatures in order to maintain a symmetrical shaft horse-power at

flight idle. If the TD control valve switch is then positioned at LOCKED, the TD valve is locked at whatever position

it is in at the time. The TD valves remain locked and the fuel correction lights remain out through all throttle

movements, unless an overtemperature condition is sensed by the amplifier. When the switch is in the AUTO or

LOCKED position, the TD valve for an engine is unlocked and set to a “take” position if TIT for the engine exceeds

approximately 1,083 _C. If a valve is unlocked by its control system to correct an overtemperature condition, the fuel

correction light for that engine comes on to indicate that the valve is unlocked.

Figure 2-6. Temperature Datum Control Valve Panel

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Note

The switches lock the TD valves only when they are positioned at

LOCKED while the throttle is in temperature-controlling range and the fuel

correction light is out.

The NULL position removes ac power from the control system amplifier; the TD valve, receiving no control signals,

returns to its null position so that it does not correct the fuel flow according to TIT. The TD valve brake is released

by dc power supplied through a FUEL & TEMPERATURE CONTROL circuit breaker on the pilot lower circuit

breaker panel.

The NULL position of these switches is used to deactivate the control systems when erratic fuel scheduling is

suspected or when the engines are not operating.

2.1.7.6 Electronic Fuel Correction Lights

The four amber press-to-test ELECTRONIC FUEL CORRECTION lights (see Figure 2-7) are located on the pilot

instrument panel. Thelights areilluminated in thetemperature-limiting range(throttles below65_)andextinguished

in thetemperature-controlling range(throttles above65_), ifthe TEMP DATUM CONTROL VALVE switches (see

Figure 2-6) are in the AUTO position. The lights will be illuminated below 65_ throttle and extinguished above 65_

throttle, but will illuminate again if temperature limit is reached when the TEMP DATUM CONTROL VALVE

switches are in the LOCKED position.

Figure 2-7. Electronic Fuel Correction Lights

2.1.7.7 Starting Control System

The starting control system (Figure 2-8) automatically controls fuel flow and ignition during ground and air starts.

Electrical power for the control circuits is supplied from the essential dc bus through engine START CONTROL

circuit breakers and the IGNITION CONTROL circuit breakers on the copilot side circuit breaker panel. The

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automatic control of the starting control system has a speed-sensitive control and a speed-sensitive valve, which is

engine-driven and performs the following functions:

1. On acceleration to 16-percent rpm — The fuel shutoff valve in the engine fuel control is opened, the ignition

relay is energized completing circuits to the ignition exciter, the engine fuel pump paralleling valve closes,

the fuel enrichment valve opens, and the manifold drip valve closes.

2. On acceleration to 65-percent rpm — Ignition system is deenergized, fuel pump paralleling valve is opened

to return pumps to series operation, manifold drip valve is deenergized (it is then held closed by pressure).

3. On acceleration to 94-percent rpm — Electronic TD control system is switched from start limiting to normal

limiting, and the speed-sensitive valve opens to allow 14th-stage bleed air to force the 5th- and 10th-stage

compressor bleed valves closed.

Figure 2-8. Engine Start Control System (Simplified Diagram)

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2.1.7.8 Normal Engine Starting Sequence

During a normal start, the following actions take place automatically (provided Chapter 8 checklist procedures have

been followed) as listed in Figure 2-9. An examination of the sequence will be helpful in understanding the overall

operation of any start.

PERCENT ENGINE RPM

(APPROXIMATE)

ACTION

CONTROLLED BY

0 to 94

TIT limited to 830_ by TD control

Speed-sensitive switch

0 to 94

5th- and 10th-stage compressor bleeds

Speed-sensitive valve

open

—

Electronic fuel correction light on

Throttle and electronic fuel correction switch

Fuel shutoff opened

Speed-sensitive switch

Fuel enrichment on

Speed-sensitive switch and fuel enrichment

switch

Fuel pumps in parallel operation

Speed-sensitive switch

16 and up

Drip valve closed

Speed-sensitive switch and pressure

Ignition on

Speed-sensitive switch

50 psig fuel manifold pressure

Fuel enrichment off

Manifold pressure switch

Starter switch released

Pilot

Fuel pumps in series operation

Speed-sensitive switch

Ignition off

Speed-sensitive switch

5th- and 10th-stage compressor bleeds

Speed-sensitive valve

closed

TIT limited by TD

Speed-sensitive switch

Figure 2-9. Normal Engine Starting Sequence

2.1.7.9 Compressor Unloading

Bleed valves are located in the 5th and 10th stages of the compressor to allow more rapid acceleration of the turbine

during starting and to minimize compressor surge or stall problems. The opening and closing of the 5th- and

10th-stage bleed valves is controlled by an engine-driven centrifugally actuated valve. On acceleration, the bleed

valves are open from 0 to 94 percent at which point the engine-driven speed-sensitive valve opens to allow control

air pressure flow from the 14th stage to close the bleed valves. The bleed valves remain closed until engine speed

decelerates below 94 percent.

2.1.7.10 ENGINE GROUND START Switches

Four spring-loaded ENGINE GROUND START switches are located on the overhead control panel (see Figure

2-10). Holding the ENGINE GROUND START switch to the START position opens the starter regulator valve to

permit bleed air from the bleed air manifold to drive the engine starter turbine. The ENGINE GROUND START

switch should be released to OFF at 60-percent engine rpm.

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Figure 2-10. Engine Starting and Fuel Enrichment Panel

2.1.7.11 Engine Fuel Enrichment Switches

The ENGINE FUEL ENRICHMENT switches are located on the engine starting panel (see Figure 2-10). They are

toggle switches with NORM and OFF positions. In NORM, each switch allows the engine fuel enrichment valve to

be controlled by the speed-sensitive control and manifold pressure switch during starting. The OFF position is

provided to permit deactivating the fuel enrichment system for any engine. During the engine starting cycle the fuel

enrichment system furnishes unmetered fuel to the TD valve to supplement normal flow through the fuel control.

This enriching starts at 16 percent rpm and lasts only until fuel manifold pressure reaches approximately 50 psi.

Note

If fuel enrichment is used and the engine does not start, the drip valve

should drain all excess fuel in the engine overboard when the start is

discontinued. If the drip valve fails to drain the excess fuel, it is advisable

to place the ENGINE FUEL ENRICHMENT switch to OFF for the next

starting attempt and motor the engine, with the condition lever in

GROUND STOP, to eliminate the accumulated fuel from the engine to

preclude a hot start or torching.

Characteristics of the starts will vary during extreme cold weather. With fuel enrichment off, light-off occurs between

22.0- and 26.5-percent rpm and several seconds may elapse between maximum turbine-starter-driven rpm and

light-off. Torching may occur also. Starting with fuel enrichment normal produces light-off between 19.0- and

27.0-percent rpm with rapid engine acceleration. If a stalled start takes place, it can be noted by rpm lag at about

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40-percent rpm and a sharp TIT increase. When theengine is still hot from previous operation, stalled starts aremore

likely to occur with fuel enrichment on. After an unsuccessful attempt to start the engine with fuel enrichment OFF,

the next attempt should be made with fuel enrichment normal.

2.1.7.12 ENGINE BLEED AIR Switches

The ENGINE BLEED AIR switches, located on the anti-icing and deicing control panel, are three-position (OFF,

ON, OVRD) toggle switches. Each switch controls a pressure-actuated, dual solenoid-controlled pressure regulator.

When the bleed-air switch is in OFF, the regulator shuts off all airflow to or from the engine. When the switch is in

the ON position, the regulator regulates airflow from the engine to the bleed-air manifold to approximately 50 psi

and prevents airflow into the engine nacelle if the bleed-air manifold pressure is above approximately 50 psi. Low

bleed-air manifold pressure will allow airflow into an engine nacelle. When the switch is in OVRD (override), the

regulator is fully open and permits air flow in either direction. It is necessary to use the OVRD position during engine

starting,nacellepreheating(ifinstalled),andforengineinletairscoopanti-icingwith theenginenotrunning. Acheck

valve is provided to prevent backflow into the engine diffuser. The bleed-air regulators receive 28-Vdc power from

the essential dc bus through the BLEED AIR FIRE SHUTOFF VALVES circuit breakers on the copilot side circuit

breaker panel. The regulators go to the closed position when deenergized.

2.1.7.13 Feather Valve and NTS Test Switch and Lights

The feather valve and NTS check system (see Figure 2-18) consist of a FEATHER VALVE AND NTS CHECK

switch, four indicator lights (one for each engine), four NTS check relays (one for each engine), and a feather valve

switch and an NTS switch in each propeller control assembly. When the FEATHER VALVE AND NTS CHECK

switch is in the VALVE position, it completes the light circuits from the essential dc bus through the lights and

contacts of each NTS check relay to the feather valve switch in each propeller control assembly. If the feather valve

is positioned by thecondition leverforfeathering thepropeller, it completes a circuit to ground for thecorresponding

indicator light. The light will come on to indicate that the feather valve is in position to feather the propeller. When

a propeller is feathered by a fire emergency control handle, the corresponding light will not come on, although the

feather valve is in the feather position. When the FEATHER VALVE AND NTS CHECK switch is in the NTS

position, it completes two circuits. One circuit is completed from the essential dc bus through each indicator light

to a set of contacts in each NTS check relay. The other circuit is completed from the essential dc bus through the coil

of each NTS check relay to the NTS check switch in the propeller control assembly. When a negative torque condition

exists, the engine NTS plunger actuates a linkage that closes the NTS switch. The NTS switch completes a circuit

to ground for the NTS check relay coil and energizes the relay. The relay actuates to provide a ground path for the

light circuit and the relay coil. The relay will remain energized, and the indicator light will glow as long as the

FEATHER VALVE AND NTS CHECK switch is in the NTS position.

2.1.8 Engine Instruments

The engine instruments are located on a panel at the center of the main instrument panel. Indicator lights for fuel

pressure warning are on the overhead control panel. For additional information on fuel flow gauges and pressure

warning lights, refer to Fuel System Indicators, paragraph 2.6.11.

2.1.8.1 Torquemeters

Each of the four torquemeters (see Figure 2-11) indicates positive and negative torque in inch-pounds. The indicated

torque is detected at the extension shaft between the engine power section and reduction gear assembly. The

torquemetersystem uses 115-volt, single-phase, acpowerfrom theacinstrument andenginefuelcontrol busthrough

the ENGINE TORQUEMETER circuit breakers on the pilot lower circuit breaker panel.

2.1.8.2 Tachometers

Each of the four tachometers (see Figure 2-11) indicates engine speed in percent of normal engine rpm. A vernier

dial on each indicator makes it possible to read to the nearest percent. The tachometer system uses a separate

engine-driven tachometer generator, mounted on each engine, that is not dependent upon the aircraft electrical system

for operation.

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2.1.8.3 Turbine Inlet Temperature Indicators

Each of the TIT indicators (see Figure 2-11) indicates temperature sensed by thermocouples in the engine turbine

inlet casing. Each indicator registers temperature in degrees Centigrade and contains a vernier scale graduated in

degrees. The indicator system uses 115-volt, single-phase, ac power from the ac instrument and engine fuel control

bus through the TURBINE INLET TEMPERATURE circuit breakers on the pilot lower circuit breaker panel.

2.1.8.4 Fuel Flow Gauges

Each of the four fuel flow gauges (see Figure 2-11) indicates flow in pounds per hour. Flow is measured at the point

where fuel enters the manifold on the engine. The indicating system receives single-phase, 115-Vac power from the

ac instrument and engine fuel control bus through the FUEL FLOW circuit breakers on the pilot lower circuit breaker

panel, and it receives 28-Vdc power from the essential dc bus through the FUEL FLOW circuit breaker on the copilot

lower circuit breaker panel.

2.1.8.5 Secondary Fuel Pump Pressure Lights

Four press-to-test secondary fuel pump pressure lights are located on the engine starting and fuel enrichment panel

(see Figure 2-10). Each light is controlled by a pressure switch on the engine fuel pump and filter assembly. The light

is normally on while the two gear pumps in the assembly are operating in parallel during engine starting (prior to

65-percent rpm). The light also illuminates at any other time if the pump paralleling valve is closed or if the primary

gear pump fails. If the light does not illuminate during starting, either the pump paralleling valve is open or the

secondary pump has failed. The lights are energized by 28-Vdc power from the essential dc bus through the SEC

PUMP IND LIGHT circuit breaker on the copilot side circuit breaker panel.

2.1.8.6 Oil Temperature Gauges

Thefouroil temperaturegauges (seeFigure 2-11)indicate oil temperature in the oil inlet lines in degrees Centigrade.

The electrical-resistance-type indicators receive 28-Vdc power from the essential dc bus through the ENGINE OIL

TEMP INDICATOR circuit breaker on the copilot side circuit breaker panel.

2.1.8.7 Oil Pressure Gauges

Four dual oil pressure gages (see Figure 2-11) register oil pressure for both the engine power sections and reduction

gears. The rear needle marked “G” on each indicator shows reduction gear oil pressure, and the front needle marked

“E” indicates power section oil pressure. The oil pressure gauges receive 26-Vac power from the instrument

transformers through engine Nos. 1, 2, 3, and 4 GEAR BOX IND OIL PRESSURE and IND ENGINE OIL

PRESSURE fuses on the pilot lower circuit breaker panel.

2.2

FIRE AND OVERHEAT DETECTION SYSTEMS

A fire detection system is provided for each engine and the APU. An overheat detection system is provided for the

engine turbine and nacelle, anti-icing, and air-conditioning. The engine and nacelle fire and overheat detection

systems are described below. Refer to the appropriate system for description of other detection systems.

2.2.1 Fire Detection System

The fire detection system for each engine and APU consists of a continuous loop detector, an amplifier, and warning

lights in the flight station. When a high temperature is detected, the amplifier unit initiates a signal to the warning

lights. These lights give a steady red glow when activated. The fire detection system receives 28-Vdc power from

the essential dc bus through the FIRE DETECTOR circuit breakers on the copilot side circuit breaker panel.

2.2.1.1 Fire Detection System Test Switch

The fire detection test switch labeled ENGINE FIRE is located on the warning system test panel (see Figure 2-12)

on theoverhead control panel. Thetest systemis providedto testtheoperationofthedetectors andthewarninglights.

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When the switch is placed in the TEST position, all five fire detection systems are activated and the warning lights

will go on. Failure of a warning light to go on indicates a break in continuity in the warning circuit.

Figure 2-12. Warning System Test Panel

2.2.1.2 Fire Detection System Warning Lights

The fire detection system warning lights consist of the master fire warning light and the lights in the fire emergency

control handles.

2.2.1.2.1 Master Fire Warning Light

A red master fire warning light and an edge-lighted panel are located on the pilot instrument panel (see Figure 2-13).

If a fire is detected by any one of the detection systems, the warning light and panel light will glow steadily. The steady

light distinguishes the signal from an overheat warning indication, which is a flashing of the same light. When the

master light indicates a fire, the lights in one of the fire emergency control handles will be on also to indicate the

location of the fire. The master fire warning light receives power from the essential dc bus through the MASTER

FIRE WARNING circuit breaker on the copilot side circuit breaker panel.

2.2.1.2.2 Fire Emergency Control Handle Lights

Each emergency control handle (see Figure 2-17) contains four indicator lights. The top two lights in the handle flash

indicating a turbine overheat, while the two lower lights glow steady when a fire is detected in the corresponding

engine.

2.2.2 Turbine Overheat Warning Systems

An overheat warning system is provided for each engine turbine. Each system consists of four thermal-switch detector

units mounted in the “hot section” of the nacelle aft of the firewall, a flasher, and indicator lights. These components

are interconnected so that an overheat condition sensed by any one of the detectors causes the lights to flash. The

detectors are connected in parallel to a loop; and if part of the detectors is inoperable, the remaining detectors can

still close the circuit to turn on the lights. A test switch permits testing all four systems at the same time. The fenwal

setting at which the detector lights will give an overheat warning is approximately 700 _F. The 28-Vdc power for

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energizing the system is supplied from the essential dc bus through OVERHEAT DETECTORS TAILPIPE circuit

breakers on the copilot side circuit breaker panel.

Figure 2-13. Master Fire Warning Light

2.2.2.1 Turbine Overheat Warning Lights

Overheat warning of the engine turbine is indicated by flashing of the master fire warning light and the lights in the

fire emergency control handles.

2.2.2.1.1 Master Fire Warning Light

A red master fire warning light and an edge-lighted panel are located on the pilot instrument panel (see Figure 2-13).

The master fire warning light and the panel light flash when any one of the engine overheat warning systems senses

an overheat condition. The lights in the fire emergency control handle for the overheated engine will flash also. The

master fire warning light receives power from the essential dc bus through the MASTER FIRE WARNING circuit

breaker on the copilot side circuit breaker panel.

2.2.2.1.2 Fire Emergency Control Handle Lights

Each fire emergency control handle (see Figure 2-17) contains two red indicator lights that flash when an overheat

condition is sensed in its respective engine.

2.2.2.2 Turbine Overheat Detector Test Switch

The overheat detection test switch labeled TURBINE OVERHEAT (see Figure 2-12) is located on the warning

system test panel on the overhead control panel. The switch has NORMAL and TEST positions. When positioned

at TEST, it closes all four of the overheat warning system circuits in the same manner as if they were closed by

detectors sensing an overheat condition. If the indicator lights all come on and flash when the switch is operated,

circuit continuity and flasher operation are satisfactory.

Note

The test switch will only check circuit continuity and that the switch is

functioning properly. Even though all indicator lights come on and flash,

this does not indicate the detectors are properly set or even operating.

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2.2.2.3 Audible Warning Test Switch

The AUDIBLE WARNING test switch (see Figure 2-12) is a three-position (TEST, NORM, SILENCE),

spring-loaded toggle switch, located on the warning system test panel. The TEST and SILENCE positions of the

switch are both test positions. Holding the switch to TEST while visually monitoring the master fire warning light

and No. 1 fire emergency control handle light provides a check of the continuity of the fire warning system. The

master fire warning light panel and the No. 1 fire emergency control handle light should illuminate with a steady

glow. When the switch is released, it should return to NORM. Holding the switch to SILENCE while simultaneously

holding the engine test switch to TEST and repeating both monitoring operations provides a check on the current

operational status of a holding relay in the system. If the master fire warning light and all fire emergency control

handle lights glow, the relay is functioning as desired. Releasing the switch from the SILENCE position should cause

it to return to NORM. Electrical power for the system is furnished from the essential dc bus through an AUDIO

WARNING circuit breaker located on the copilot side circuit breaker panel.

Note

There is no audible fire warning on any Navy C-130T aircraft. This switch

only tests a holding relay.

2.2.3 Nacelle Overheat Warning Systems

An overheat warning system is provided for each nacelle. Each system consists of eight thermal-switch detector units,

seven mounted in the nacelle area forward of the firewall, one mounted in the horse collar area, and a warning light

on the copilot instrument panel. A test switch is provided for testing all four warning systems simultaneously. The

purpose of each system is to warn of an overheat condition in the area around the engine compressor section. Overheat

in this areacan result from thenacellepreheat valvebeing openedoraruptureoccurringin thebleed-airsystemducts.

Theoverheatconditioncould alsoresult fromfire. Theoverheat conditioncan bedetected byany oneofthedetectors,

which are connected in parallel to a loop. The fenwal setting at which the detector lights will give an overheat warning

is approximately 300 _F.

2.2.3.1 Indicator Lights

Fournumberednacelleoverheatwarninglightsandatestpanel(seeFigure2-14)arelocated onthecopilotinstrument

panel. If overheat is detected in any nacelle, the corresponding overheat warning light and the light on the test panel

will go on. Power for the overheat warning lights comes from the essential dc bus through the OVERHEAT

DETECTORS NACELLE TAILPIPE circuit breakers on the copilot side circuit breaker panel. The indicator lights

receive 28-Vdc power from the essential dc bus through the NACELLE OVERHEAT NAMEPLATE and

OVERHEAT DETECTORS NACELLE circuit breakers on the copilot side circuit breaker panel.

2.2.3.2 Nacelle Overheat Test Switch

A nacelle overheat test switch is located on the copilot instrument panel next to the warning lights (see Figure 2-14).

Operation of the test switch closes all four nacelle overheat warning circuits simultaneously, causing all four warning

lights and the panel lights to glow as long as the switch is held in TEST. Failure of a light to go on indicates a break

in continuity in the warning circuit.

Note

The test switch will check only circuit continuity and that the switch is

functioning properly. Even though all indicator lights go on, this does not

indicate that the detectors are properly set or even that they are operating.

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Figure 2-14. Nacelle Overheat Control Panel and Lights

2.2.4 Bleed Air Duct Overheat Detection System (ODS)

An overheat warning system is provided for the aircraft’s high-pressure bleed air ducts. The system will provide

warnings upon detection of small leaks in the bleed air duct system, which if left undetected could cause extensive

collateral damage to other aircraft systems and structures. The system consists of a control unit, a relay panel, an ODS

control panel, and several continuous discrete sensing elements divided into 11 zones. The sensing elements will

detect a rise in temperature and are set to provide visual and audible warnings at temperatures of 310 _F for elements

located outsidetheaircraft and 255 _F forelements withinthefuselagecompartment (seeFigure2-15).Ifanoverheat

condition is detected, warning lights in the flight station will illuminate and the landing gear warning horn will pulse

at 90 cycles per minute. The control unit in the cargo compartment will also display a code that can be used for further

troubleshooting. When the overheat condition diminishes, the system will reset and the sensing elements return to

theiroriginal condition with no external actionrequired. Abroken sensingelement willcontinuetofunction oneither

side of the break; however, the system will not properly test. The system receives 115-Vac power from the essential

ac bus through the LHS SENSORS and RHS SENSORS circuit breakers and 28-Vdc power from the essential dc

bus through the ODS ALARM and ODS CONTROL circuit breakers located on the pilot’s lower circuit breaker

panel.

2.2.4.1 ODS Control Panel

The ODS control panel is located on the flight engineer’s overhead panel (Figure 2-15). The control panel provides

visual indications of an ODS warning and incorporates a warning horn silence button and system test switch. Eleven

NVIS red indicator lights correspond to the 11 zones of the sensor elements and are grouped into six output alarms.

Alarm signal 1 illuminates all three indicator lights in the left wing. Alarm signal 2 illuminates all three indicator

lightsintherightwing.Alarmsignal3illuminatesboththetransfuselageandforwardfuselageindicatorlights.Alarm

signal 4 illuminates the aft fuselageindicator light. Alarm signal 5 illuminates theAPU indicatorlight. Alarm signal

6 illuminates the air conditioning indicator light.

2.2.4.1.1 ODS Test Switch

A lever locked NORMAL/TEST toggle switch located on the ODS control panel allows the operator to activate a

system self-test. The switch is spring-loaded to the NORMAL position. Placing the switch to the test position will

illuminate the 11 indicator lights, the ODS master warning light, and cause the landing gear warning horn to sound

in pulsing mode (90 cycles per minute).

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2.2.4.1.2 ODS Horn Silence Switch

A HORN SILENCE switch is located on the ODS control panel and allows the ODS warning horn to be silenced

during an alarm condition or self-test.

2.2.4.2 ODS Master Warning Light

An amber ODS master warning light is located on the lower left side of the engine instrument panel (Figure 2-11).

If an overheat condition is detected, the master warning light will illuminate steady along with the pulsing landing

gear warning horn and corresponding alarm indicator lights on the overhead control panel.

2.2.4.3 ODS Control Unit

The ODS control unit is located on the starboard side of the cargo compartment. The control unit monitors sensing

elements mounted in 11 zones throughout the aircraft and groups them into 6 alarm zones. The control unit’s

monitoringcircuitsrecordanyalarmconditionand transmitsignals tothecontrolpanel intheflightstation toindicate

where the overheat condition exists. Specific alarm and malfunction codes are displayed on LED readouts on the face

panel of the control unit. All malfunction and alarm codes are stored in memory for maintenance and repair purposes.

2.3

FIRE EXTINGUISHING SYSTEM

A two-shot bromotrifluoromethane (BT) fire extinguishing system (see Figure 2-16) is connected through a series

of directional-flow valves to each of the four engine nacelles and to the APU compartment. The extinguishing agent

is contained in two bottles mounted in the left wheelwell. Each bottle contains approximately 27 pounds of agent.

One bottle is discharged each time the system is actuated. A check valve prevents a discharged bottle from being

recharged when a fresh bottle is fired. Each bottle is charged to approximately 600 psi with nitrogen, the nitrogen

acting as a propellant for the BT. Individual pressure gauges on each bottle show charged pressure.

Repeated or prolonged exposure to high concentrations of BT or

decomposition products should be avoided. BT is a narcotic agent of

moderate intensity but of prolonged duration. It is considered to be less

toxic than carbon tetrachloride, methyl bromide, or the usual products of

combustion. BT is safer to use than these fire extinguishing agents.

However, normal precautions should be taken, including the use of oxygen

when available.

2.3.1 Fire Extinguishing System Controls

The fire extinguishing system controls are located on the fire emergency control panel (see Figure 2-17) forward of

the overhead electrical control panel. The fire extinguishing system control circuits use dc power supplied from the

battery bus through three FIRE EXT circuit breakers on the pilot side circuit breaker panel. One circuit breaker labeled

FIRE EXT is in the control circuit and the other two labeled FIRE EXT NO. 1 and NO. 2 are in the squib firing circuits

(one for each squib).

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Figure 2-17. Fire Emergency Control Panel

2.3.1.1 AGENT DISCHARGE Switch

A three-position (NO. 1, OFF, NO. 2) toggle switch located on the fire emergency control panel (see Figure 2-17)

controls the discharge of the bottles. The AGENT DISCHARGE switch is spring-loaded to the OFF position. The

agent will not discharge unless a fire emergency control handle is pulled. The fire emergency handle circuit powers

the correct sequence of solenoid directional control valves in the system to direct flow ofagent to the selected engine

when one of the bottles is fired. The directional flow valves move in the same order as the handles are pulled. If two

fire emergency control handles are pulled, the agent will be routed to the engine for the last handle pulled. In order

to route agent to the engine for the first handle pulled, the first handle must be pushed in and pulled again.

2.3.1.2 Fire Emergency Control Handles

The five plastic fire emergency control handles are mounted on the fire emergency control panel (see Figure 2-17).

They operate emergency shutdown switches for the APU and the four engines. When an engine handle is pulled out,

it closes dc circuits to operate valves that isolate the engine as follows:

1. The shutoff valve on the engine fuel control is closed.

2. The engine oil shutoff valve is closed.

3. The firewall fuel shutoff valve is closed.

4. The firewall hydraulic shutoff valves are closed.

5. The engine bleed-air regulator is closed.

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6. Engine starting control circuits are deenergized.

7. The propeller is feathered.

8. The fire extinguisher system directional flow valves are positioned for routing agent to the engine. The

extinguishing AGENT DISCHARGE switch is armed.

When the APU handle is pulled, the APU is isolated as follows:

1. The fuel shutoff valve is closed.

2. The bleed-air valve is closed.

3. The fire extinguisher system directional flow valve is positioned for routing agent to the APU.

4. The extinguisher AGENT DISCHARGE switch is armed. The APU door closes.

2.4

PROPELLERS

Each engine is equipped with a Hamilton Standard, four-blade, electro-hydromatic, full-feathering, reversible-pitch

propeller. The propeller operates as a controllable-pitch propeller for throttle settings below flight idle and as a

constant-speed propeller for throttle settings of flight idle or above. The major components of the propeller system

are the propeller assembly, the synchrophasing system, the control system, and the anti-icing and deicing system.

2.4.1 Propeller Assembly

The propeller assembly is made up of the propeller blades, barrel assembly, dome assembly, spinner assembly,

control assembly, and the anti-icing and deicing assembly.

2.4.1.1 Propeller Blades

The propeller blades are solid aluminum alloy with hollow shanks for weight reduction. On the mounting end of the

blades are located the blade gear segments, thrust bearings, oil seals, and deicing rings.

2.4.1.2 Barrel Assembly

The principal functions of the barrel assembly are to retain the blades within the propeller assembly, to provide the

necessary means of attaching the propeller to the engine shaft, and to transmit engine torque to the blades. The barrel

assembly is made in two sections that are bolted together to retain the propeller blades. An extension on the rear half

of the assembly is machined to fit over the splined engine shaft. The pitchlock assembly, located inside the barrel

assembly, consists of a stationary pitchlock ratchet ring, a pitchlock rotating ratchet, a pitchlock valve, and fly-weight

assembly. The pitchlock prevents the blades from decreasing pitch if hydraulic pressure is lost, and it locks the blades

at a fixed pitch angle if overspeeding occurs. The stationary and rotating pitchlock ratchet rings are held disengaged

by propeller oil pressure and are spring loaded to engage when pressure is lost. When the ratchet rings are engaged,

the propeller can increase pitch to allow feathering. In an overspeed condition, the flyweight assembly will remove

oil pressure and allow the pitchlock to engage and prevent a decrease in blade angle. When the overspeed condition

is corrected, oil pressure is restored to release the pitchlock.

2.4.1.3 Dome Assembly

The dome assembly is mounted on the forward section of the barrel assembly. It contains the pitch changing

mechanism and the low-pitch stop assembly. The pitch changing mechanism converts hydraulic pressure into

mechanical torque. Its main parts are a piston assembly, a stationary cam, a rotating cam, and the dome shell. The

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piston is a double-walled assembly that fits over the two cams and inside the dome shell. The piston is held in place

by rollers that ride in the cam tracks of both cams. The rear of the rotating cam is connected by beveled gears to the

propeller blades. As hydraulic pressure is applied to the piston, causing it to move, the rollers riding in the cam tracks

turn the rotating cam, changing the blade angle. The low-pitch stop in the dome mechanically stops the piston from

decreasing blade angle below the flight range. The low-pitch stop is retracted to allow lower blade angles during

ground operation.

2.4.1.4 Control Assembly

The propeller control assembly is mounted on the aft extension of the propeller barrel but does not rotate. It contains

the oil reservoir, pumps, valves, and control components that supply the pitch changing mechanism with hydraulic

pressure of the proper magnitude and direction to vary the propeller blade angle as required for the selected operating

condition. The valve housing assembly section of the control assembly contains the flyweight speed-sensing pilot

valve, feather valve, feather solenoid valve, and feather actuating valve. The pump housing assembly contains a

scavenge, main, and standby pump, and an electrically driven, double-element, auxiliary pump. The flow of fluid

from these pumps is controlled by the valves in the valve housing assembly to accomplish the desired propeller

operation. All mechanical and electrical connections necessary for propeller operation are made through the control

assembly.

The mechanical connections are linkages from the engine control system and the NTS system. The electrical

connections are for oil level indication, pulse generator coil, auxiliary pump motor, synchrophasing system,

anti-icing and deicing systems, and the electrical feathering system.

2.4.1.5 Spinner Assembly

Thespinnerassembly,whichimprovestheaerodynamiccharacteristicsofthepropellerassembly,encloses thedome,

barrel, and control assemblies. It consists of a front section, a rear section, and a nonrotating afterbody assembly.

Cooling air is admitted through an air inlet at the front of the spinner and passes over the dome assembly, barrel

assembly, and control assembly fins, and exhausts through vents in the engine nacelle.

2.4.1.6 Anti-Icing and Deicing Assembly

The anti-icing and deicing assembly is made up of stationary and moveable contacts for conducting electrical power

to the resistance-type heating elements on the leading edge and shank of each blade and the entire spinner assembly.

Continuous anti-icing heaters cover the spinner front section and the afterbody assembly. Cyclic deicing heaters

cover the remainder of the spinner front section, the spinner rear section, the blade shanks, and part of each blade’s

leading edge. The contact ring holder assembly mounted on the aft end of the barrel assembly contains four sliprings

and four contact brush housing assemblies. Power from the aircraft electrical system is transmitted through the

brushes and sliprings to the spinner anti-icing and deicing elements and the blade’s deicing elements.

2.4.1.7 Propeller Low Oil Level Warning Lights

A PROPELLER LOW OIL WARNING light for each propeller is located on the copilot side shelf(see Figure2-18);

a PROP LOW OIL quantity light, which acts as a master warning light, is located on the engine instrument panel

(see Figure 2-11). The propeller low oil warning system is controlled by a float-actuated switch in each propeller

control assembly. When the oil quantity for each propeller drops approximately 2 quarts below normal, the

float-actuated switch closes and illuminates the PROPELLER LOW OIL WARNING light for that engine and the

PROPLOWOILquantity light.Iftheoil quantitybecomes lowforanotherpropeller, theonly indicationwill befrom

the PROPELLER LOW OIL WARNING light for that engine. The low oil warning lights receive 28-Vdc power from

the essential dc bus through the PROP LOW OIL LEVEL circuit breaker on the copilot side circuit breaker panel.

2.4.2 Propeller Speed Control System

The speed of the propeller is controlled by the propeller governing system and the synchrophasing system.

ORIGINAL

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Figure 2-18. Propeller Controls

2.4.2.1 Propeller Governing System

The principal function of the propeller governing system is to maintain constant engine operating rpm. Propeller

governing is accomplished by the action of the flyweight speed-sensing pilot valve. This valve is controlled by the

mechanical action of the flyweight opposing the tension of the speeder spring. When the propeller is in an on-speed

condition, the pilot valve meters sufficient fluid to the increased pitch or forward side of the dome assembly piston

to overcome twisting movement and maintain the required blade angle. When an overspeed condition occurs, the

flyweight force overcomes the speeder spring force, and the pilot valve moves to increase the flow to the

increased-pitch side of the piston. If the propeller slows below governed speed, the force of the speeder spring

overcomes the force exerted by the flyweights, and the pilot valve meters fluid to the aft side of the dome assembly

piston to decrease blade angle and allow the propeller to increase speed. The low-pitch stop prevents the propellers

from decreasing blade angle below the flight range while the throttles are in the flight range.

2.4.2.2 Synchrophasing System/Electronic Governing

The synchrophaser electronic unit provides circuits for the following governing functions: speed stabilization

(derivative), throttle anticipation, and synchrophasing. The propeller mechanical governor will hold a constant speed

in theflight rangebut throttlechanges willcausethegovernortooverspeed orunderspeed whiletrying tocompensate

for the change in power. A stabilization circuit stabilizes the mechanical governor during these changes when the

propeller governor control switch is in the NORMAL/NORM position by sending a signal to the speed bias servo

control motor to change the speeder spring compression. The throttle anticipation circuit stabilizes the propeller

speed during rapid movement ofthethrottlewhen thepropellergovernorcontrol switch is in the NORMAL position.

Throttle movement rotates the anticipation potentiometer in the propeller control assembly sending a signal to the

2-33

ORIGINAL

01-75GAL-1

anticipation circuit which sends an amplified signal to the speed bias servo control motor to change the speeder spring

compression. The synchrophasing system acts to keep all the propellers turning at the same speed, and it maintains

a constant blade rotational position relationship to decrease vibration and to lower the noise level. The system uses

either No. 2 or No. 3 engine as the master engine, and it relates the blade position of the other three engines to the

master. The blade position of a slave engine is changed by moving the pilot valve to increase or decrease the speed

of that engine. The synchrophasing circuit determines the blade position by comparing an electrical pulse generated

by each slave propeller to a modified pulse from the master propeller. If the blades are in the correct position, the

resultant voltage of the slave and master pulse will be zero. Any deviation in blade position will produce a positive

or negative voltage from the two compared pulses. This voltage drives the speed bias servo control motor to change

the speeder spring compression, correcting the blade position. If propeller operation is erratic, see paragraph 11.4.2,

Propeller Malfunctions. The electronic propeller governing system is powered by the essential dc bus through the

SYNCHROPHASER circuit breaker on the copilot side circuit breaker panel.

2.4.3 Negative Torque Signal Lockout System

The propeller is equipped with a lockout system to deactivate the NTS system for throttle settings below FLIGHT

IDLE. When the throttle is moved below FLIGHT IDLE, a cam moves the actuator away from the NTS plunger and

renders the system inoperative. This is necessary to prevent a propeller from receiving a possible negative torque

signal at high landing speeds when the throttles are moved toward reverse, with resultant asymmetrical power

problems.

2.4.4 Propeller Controls

Propeller controls include the throttles, condition levers, SYNCHROPHASE MASTER switch, PROP RESYN-

CHROPHASE switch, PROPELLER GOVERNOR CONTROL switches, fuel governor check switches, and feather

override buttons.

2.4.4.1 Throttles

Each throttle (see Figure 2-4) is mechanically linked through the engine coordinator to an input shaft on the propeller

control assembly. When the throttle is in the governing range, between FLIGHT IDLE and TAKE-OFF positions,

the input shaft rotates with throttle movement but has no effect on propeller speed. When the throttle is in the range

below FLIGHT IDLE, any movement of the throttle is transmitted to the speed-sensing pilot valve to increase or

decrease blade angle. The maximum negative blade angle is obtained when the throttle is at MAXIMUM REVERSE.

Approximate minimum thrust angle is obtained when the throttle is at GROUND IDLE. When the throttle is moved

below FLIGHT IDLE, a cam locks out the NTS system and a switch interrupts synchrophaser signals to the

propeller.

2.4.4.2 Engine Condition Levers

The engine condition levers (see Figure 2-4) serve primarily as feathering and unfeathering controls. Each lever is

mechanically linked to the engine coordinator, which transmits the motion of the lever to the propeller linkage only

when it is moved to the FEATHER position. When pulled to FEATHER, the condition lever also actuates switches

to turn on theelectrically-driven auxiliary pump in thepropellercontrol assembly and thepropellerblades aremoved

to the feather angle. For unfeathering, the engine condition lever is held in the AIRSTART position. A switch is

actuated to turn on the propeller auxiliary pump, and the pump continues to operate as long as the lever is held in

this position. When the engine condition lever is in the AIRSTART position and the auxiliary pump is operating,

fluid is routed to the aft side of the dome assembly piston to move the blades to low-pitch angle. When the condition

lever is in GROUND STOP or RUN positions, the propeller is controlled normally and the lever has no effect on its

operation.

ORIGINAL

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01-75GAL-1

2.4.4.3 SYNCHROPHASE MASTER Switch

The SYNCHROPHASE MASTER switch (see Figure 2-19) is located on the flight control pedestal. This

three-position (ENG 2, OFF, ENG 3) toggle switch controls the operation of the synchrophase system and selects

the engine to be used as the master. When the switch is in the ENG 2 position, the No. 2 engine is selected as the

master and the other propeller phase angles are referenced to this engine. When the switch is in the OFF position and

the propeller governor control switches are in the NORMAL position, the synchrophasing system is turned off and

the propellers operate in normal governing. If the propeller governor control switches are in the MECH GOV

position, the propellers operate in mechanical governing. When theswitch is in theENG 3 position, theNo. 3 engine

is the master and the other propeller phase angles are referenced to this engine.

2.4.4.4 Prop Resynchrophase Switch

The prop resynchrophase switch (see Figure 2-19) is a two-position (NORMAL, RESYNC) toggle switch located

on the flight control pedestal. The switch is spring-loaded and shear-wired to the NORMAL position. The prop

resynchrophase switch shall not be positioned to RESYNC except when performing the propeller re-indexing

procedure and should be used only for correcting an out-of-sync or off-speed condition by performing the complete

re-indexing procedure. These airplanes will have a decal installed above the prop resynchrophase switch on the flight

control pedestal stating SOLID STATE SYNCHROPHASER INSTALLED. REFER TO APPROVED FLIGHT

MANUAL FOR OPERATION. Power is supplied through the SYNCHROPHASER 28V essential dc circuit breaker

located on the copilot side circuit breaker panel.

CAUTION

The PROP RESYNCHROPHASE switch shall not be positioned to

RESYNC except when performing the propeller reindexing procedure and

should be used only for correcting an out-of-sync or off-speed condition by

performing the complete reindexing procedure.

2.4.4.5 Propeller Governor Control Switches

The four PROPELLER GOVERNOR CONTROL switches are two-position (NORMAL, MECH GOV) guarded

toggle switches located on the copilot side shelf (see Figure 2-18). When the switches are in the NORMAL position,

the throttle anticipation and speed stabilization (derivative) circuits are operative, and if the SYNCHROPHASE

MASTER switch is positioned to either master engine, the blade rotational position of the slave engines is related

to the master by the synchrophasing system. Placing a switch in the MECH GOV position disconnects the electrical

speed control to that propeller and the speed of the propeller is controlled by basic mechanical governing.

2.4.4.6 Fuel Governing Check Switches

The four FUEL GOVERNING check switches located on the aft end of the overhead control panel, provide a means

of checking the operation of the propeller pitch-lock mechanism and the engine fuel control governor.

CAUTION

The fuel governing check switches must never be used in flight.

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01-75GAL-1

Figure 2-19. Synchrophaser Control Switch Panel

2.4.4.7 Feather Override Buttons

Four FEATHER OVERRIDE buttons are located on the copilot side shelf (see Figure 2-18) to provide a means of

manually stopping the propeller auxiliary pump. When a condition lever is moved to FEATHER, electrical circuits

are completed to the feather solenoid valve, propeller auxiliary pump, pressure cutout switch, and a holding circuit

to thefeather overridebutton. Ifa FEATHER OVERRIDE button is manually pulled aftera condition lever is placed

in FEATHER, the electrical circuits for propeller feathering are interrupted. If a FEATHER OVERRIDE button does

not pop out after the feather cycle is complete, it should be pulled manually to prevent damage to the propeller

auxiliary pump. The holding coil of the override button receives 28-Vdc power from the essential dc bus through the

FEATHER & AIRSTART or EMER FEATHER circuit breaker on the copilot side circuit breaker panel.

2.5

OIL SYSTEM

Independent oil systems, one for each engine, supply lubrication to the engine power section and the reduction gear

assembly. An oil tank with a 12-gallon usable oil capacity is located in each nacelle above the engine. The oil tanks

are provided with a pendulum-tube oil pickup to prevent loss of oil pressure under negative-g conditions. The

pendulum will accompany the oil in a negative-g condition, assuring positive oil pressure. Oil from the tank enters

the power section and the reduction gear assembly, where it is circulated and returned by scavenger pumps through

a heat exchanger and oil cooler back into the oil tank. Hot oil passing through the heat exchanger heats the engine

fuel and prevents ice from forming in the fuel filter. Air flowing through an oil cooler duct and over the coils of the

oil cooler absorbs excess heat from the oil. A bleed-air ejector, located in the oil cooler air exit duct, provides

additional airflow through the oil cooler for ground operations. A thermostatic element in the oil tank return line

controls the oil temperature by regulating the amount of air flowing through the oil cooler duct. Four motor-operated

valves provide an emergency means of shutting off oil flow to the engines when the fire emergency control handles

are pulled. Oil used in the aircraft must conform to the specification and grade listed in the servicing diagram (see

Figure 3-1).

2.5.1 Oil System Controls

Oil system controls availableto theflightcrew aretheoil coolerflap switches and thefireemergency control handles.

ORIGINAL

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01-75GAL-1

2.5.1.1 Oil Cooler Flap Switches

Airflowthroughtheoilcoolerisgoverned byacontrollableoil coolerflap whichrestricts theopening oftheoilcooler

air exit duct, and by an oil cooler augmentation valve which controls the flow of fourteenth stage compressor bleed

air to an ejector assembly in the oil cooler exit duct. Four, four-position (AUTOMATIC, OPEN, CLOSE, FIXED)

oil cooling toggle switches are located on the oil cooling control panel on the overhead control panel (see Figure

2-20). These switches control the electrical circuits of the oil cooler flap actuators and the oil cooler augmentation

valves. In the AUTOMATIC position, the oil cooler flap position is regulated by a thermostatic unit to maintain an

oiltemperatureofapproximately80 _C.IntheOPENorCLOSEpositions(spring-loaded)thethermostat isexcluded

from the circuit and the actuator is directly energized to open or close the oil cooler flap. When the switch is moved

to FIXED, the actuator is deenergized and the flap will remain in the position it was in prior to moving the switch.

TheOPEN, CLOSE, and FIXED positions areused to control theoil coolerflapactuatormanuallyifthethermostatic

control unit fails. The oil cooler flap actuators receive 28-volt, dc power from the essential dc bus through OIL

COOLER FLAPS circuit breakers on the copilot side circuit breaker panel.

2.5.1.2 Oil Cooler Augmentation Switches

Oil coolerairflow augmentation is provided by ahigh-pressure bleed-airejector located in theoil coolerair exit duct.

Four two-position (ON, OFF) toggle switches on the overhead control panel operate oil cooler augmentation valves

that control the flow of 14th-stage compressor bleed air to the ejectors (see Figure 2-21). The oil cooler flap must

be at least 90-percent open, the throttle must be in the ground operating range, and the ENGINE GROUND START

switch must be in the OFF position before power is available to the OIL COOLER AUGMENTATION switch. The

oil cooler augmentation valves are energized through the OIL COOLER AUGMENTATION switches by 28-Vdc

power from the essential dc bus, through the START CONTROL circuit breakers on the copilot side circuit breaker

panel.

2.5.1.3 Fire Emergency Control Handles

Motor-operated shutoff valves, energized by 28-volt essential dc powerand controlled by thefire emergency control

handles (see Figure 2-17), are installed in the engine oil systems to shut off the flow of oil to the engine during an

emergency. See paragraph 2.3.1.2 for other functions of the fire emergency control handles.

Figure 2-20. Oil Cooler Flap Control Panel

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ORIGINAL

01-75GAL-1

Figure 2-21. Oil Cooler Augmentation Control Panel

2.5.2 Oil System Indicators

The oil system indicators are an oil quantity gauge for each engine, a low-oil-quantity warning light, and an oil cooler

flap position indicator for each engine. All of the indicators are located on the engine instrument panel (see

Figure 2-11).

2.5.2.1 OIL QUANTITY Gauges

Four OIL QUANTITY gauges (see Figure 2-11), one for each engine oil system, are located on the engine instrument

panel. Each indicator is calibrated from 0 to 1 (empty) to F (full) in increments of 1 gallon. The indicators are

energized by 28-Vdc power from the essential dc bus through the OIL QUANTITY INDICATOR circuit breakers

on the copilot side circuit breaker panel.

2.5.2.2 Low Oil Quantity Warning Light

A low-oil-quantity warning light (see Figure 2-11) is located on the engine instrument panel. A microswitch in each

engine oil tank quantity transmitter, actuated by the tank float arm, will cause the light to illuminate when any oil

tank level drops to 4 gallons. Operation of the light is independent of quantity indicators; however, no additional

warning is supplied should a second tank reach a low level. The warning light receives 28-Vdc power from the

essential dc bus through the engine oil quantity light circuit breaker on the copilot side circuit breaker panel.

2.5.2.3 Oil Cooler Augmentation Valve Position Indicators

Fouroil cooleraugmentation valveposition indicators, oneforeach valve, arelocatedon theengineinstrumentpanel

(see Figure 2-11). The press-to-test indicators are connected electrically to pressure switches in the ejector bleed-air

duct downstream of the augmentation valves and illuminate (OPEN) when the pressure switch closes. The lights are

energized by 28-Vdc power from the essential dc bus through the OIL COOLER VALVE OPEN IND circuit breaker

on the copilot lower circuit breaker panel.

ORIGINAL

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01-75GAL-1

2.5.2.4 OIL COOLER FLAP Position Indicators

Four OIL COOLER FLAP position indicators (see Figure 2-11), one for each engine oil system, are located on the

engine instrument panel. The indicators are electrically connected to position transmitters geared to the oil cooler

flap actuators. The indicator dials, calibrated from 0 to OPEN in increments of 10 percent, indicate the percentage

that the oil cooler flaps have opened. The indicators are energized by 28-Vdc power from the essential dc bus through

the OIL COOLER FLAPS circuit breakers on the copilot side circuit breaker panel.

2.6

FUEL SYSTEMS

Note

On aircraft incorporating AFC-424, all of the fuel tanks are equipped with

baffle systems fabricated from reticulated, explosion-fire suppressant

conductive foam.

Thefuel system is amodified manifold-flow-type,incorporating afuel crossfeedsystem, asingle-point refuelingand

defueling system, and a fuel dump system. The system provides fuel supply for the four engines and the auxiliary

power unit. It is adaptable to a number of flow arrangements (see FO-1). For systems diagrams, see FO-1. Fuel

specifications and grades are listed in Figure 3-1. Total usable capacity of the fuel tanks is shown in Figure 2-22.

For system management, refer to paragraph 8.28. Aircraft limitations resulting from use of emergency fuels are

discussed in Chapter 4.

2.6.1 Fuel Flow

Each engine may be supplied fuel either directly from the main respective fuel tank or through the crossfeed manifold

system from any tank. Fuel for the APU is routed directly from the No. 2 fuel tank.

2.6.2 Refueling and Defueling

All wing tanks and the fuselagetank (ifinstalled) may be refueled or defueled from thesingle-point ground refueling

and defueling receptacle located in the right main landing gear fairing. When refueling, fuel is routed from the

single-point receptacle through the refueling manifold to each tank through separate supply lines from the manifold.

Fuel level in the tanks is controlled by float shutoff valves. All tanks except No. 1 and 4 contain one shutoff valve

that allows the tanks to be filled to their maximum capacity, maintaining a minimum of 3-percent expansion space.

Two shutoff valves in both the No. 1 and 4 tanks are located at different levels. These valves are controlled by the

refueling pod disconnects.

The low-level valves operate when the refueling pods are installed and allow the tanks to become only partially full.

When the pods are removed, the high-level valves operate and allow the tanks to fill completely. Refueling and

defueling are controlled at the single-point refueling control panel located in the right landing gear fairing. As an

alternate method of refueling, the main and external tanks may be fueled separately through a filler opening in the

top of each tank. The fuselage tank does not incorporate a manual filler opening, but it can be refueled by transferring

fuel from the wing tanks, or from the single-point refueling control panel, and selecting the OPEN position of the

fuselage tank fill switch on the auxiliary fuel control panel. In defueling the wing tanks, fuel flows into the crossfeed

manifold through the refueling manifold, and overboard through the single-point refueling receptacle. Defueling flow

from the fuselage tank is into the refueling manifold and overboard through the single-point refueling receptacle.

2.6.3 Internal Tanks

Six fuel tanks are located within the wing (see FO-1). The No. 1, 2, 3, and 4 tanks are integral and use sealed wing

structure for tank walls. The left and right auxiliary fuel tanks are each comprised of units of three bladder cells. The

three cells are interconnected to form one assembly and are laced within the center wing section. Each of the six tanks

has a three-phase, ac-powered boost pump to ensure fuel flow. The water removal system, located in each main tank,

maintains the fuel level around the boost pump when the aircraft is in a nosedown attitude with low fuel level in the

tank. The No. 1, 2, 3, and 4 tanks have, in addition, a dump pump that is used for fuel dumping and transfer.

2-39

ORIGINAL

01-75GAL-1

USABLE FUEL

UNUSABLE FUEL (JP-4)

BOOST PUMP ON

BOOST PUMP OFF

GAUGE

TANK

GAL

JP-4

JP-5/JP-8

GAL

READING

GAL.

READING

No. 1

1,288

8,372

8,758

475

397

No. 2

1,186

7,709

8,065

592

501

Left Ext

1,379

8,964

9,377

137

Left Aux

901

5,857

6,127

Right Ext

1,379

8,964

9,377

137

Right Aux

901

5,857

6,127

No. 3

1,186

7,709

8,065

592

501

No. 4

1,288

8,372

8,758

475

397

Lines

TOTALS

Main Tanks Only

4,948

32,162

33,646

377

Main and Aux Tanks

6,750

43,876

45,900

495

Main and Ext Tanks 3

7,706

50,090

52,400

651

100

Main, Aux, and Ext Tanks

9,508

61,804

64,654

769

118

Notes

The external tanks are equipped with dual pumps and a single pump off will not affect the unusable fuel quantity in

these tanks.

The unusable fuel for an auxiliary tank is the quantity in the tank at the time the boost pump becomes inoperative.

External tank installation (Lear Siegler part no. 305J001).

Fuel quantities based on 6.5 pounds per U.S. gallon for JP-4 and 6.8 pounds per U.S. gallon for JP-5.

Pounds (LB) and gallons (Gal) are rounded off to the nearest whole number.

Figure 2-22. Fuel Quantity Data Table (Without AFC-424 Foam) (Sheet 1 of 2)

2.6.4 Water Removal System

The water removal system provides continual water removal from the tank low points during boost pump operation.

The system consists of two ejectors, a check valve, a strainer, and associated plumbing in each main tank. The ejectors

are connected by plumbing to the boost pump discharge line, and a part of the boost pump fuel flow is routed through

each ejector housing and discharged through its nozzle. This fuel flow through the ejectors causes a differential

pressure, and additional fuel is drawn from between the lower wing panel risers and is ejected into the surge box.

Anytime the fuel boost pump is operating, the fuel will be continually stirred to prevent water from settling in the

bottom of the tank.

2.6.5 Vent System

All of the six wing fuel tanks are vented to the atmosphere to equalize pressure at all times. Tank Nos. 2 and 3, and

the left and right auxiliary tanks have a wraparound vent system. The wraparound system permits venting for these

tanks even though the aircraft is not in a wings-level attitude. The outboard tanks (Nos. 1 and 4) are vented by a

float-controlled vent valve to prevent fuel loss overboard on the ground when the aircraft is not in a wings-level

attitude and in flight when the wings deflect upward. Vent air leaving the tank passes through a drainbox on its way

overboard. Any fuel entering the vent lines because of an aircraft change of attitude collects in the drainbox and is

returned to the tank continuously by the water removal system in the inboard and outboard tanks and by the jet pump

eductors in the auxiliary tanks. Boost pump pressure is necessary for the water removal system and jet pump eductors

ORIGINAL

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01-75GAL-1

to operate. The No. 1 and 4 tanks also contain a pressure-relief valve that vents tank pressure directly to the jettison

manifold, downstream of the dump mast shutoff valve.

The external tanks are vented through the spaces at the top of the bulkheads separating the tank compartments and

through the fuel vent line. The vent line runs from the forward compartment of each tank through the pylon and up

into the wing trailing edge, where it vents to the atmosphere. Fuel will not fill the vent line because each tank is

separated by compartments. The vent line is at the top of the tanks and runs upward to the wing.

USABLE FUEL -- BOOST PUMP ON

UNUSABLE FUEL -- BOOST PUMP OFF

TANK

GAL

GAUGE READING

JP--5/JP--8

No. 1

1,222

8,310

390

No. 2

1,126

7,657

490

LEFT EXT

1,309

8,901

LEFT AUX

855

5,814

RIGHT EXT

1,309

8,901

RIGHT AUX

855

5,814

No. 3

1,126

7,657

490

No. 4

1,222

8,310

390

TOTALS

Notes

MAIN TANKS ONLY

4,696

31,934

The external tanks are equipped with dual

pumps. A single pump off will not affect the unus-

MAIN AND AUX

able fuel quantity in these tanks.

6,404

43,562

TANKS

The unusable fuel fo an auxiliary tank is the

quantity in the tank at the time the boost pump be-

MAIN AND EXT

comes inoperative.

7,314

49,736

TANKS

Fuel quantities based on 6.8 pounds per US

gallon for JP--5/JP--8. Pounds per gallon are

MAIN, AUX AND

rounded off to the nearest whole calculated number.

9,024

61,364

EXT TANKS

For gauge readings, round down to the nearest 10

pounds.

To calculate weights for JP--4 fuel, multiply fuel

quantities by 6.5 pounds per gallon.

Figure 2-22. Fuel Quantity Data Table (With AFC-424 Foam) (Sheet 2)

2.6.6 Crossfeed Primer System

A press-to-actuate CROSSFEED PRIMER VALVE button is located on the fuel control panel (see Figure 2-23). This

button, when pressed, moves the motor-operated crossfeed fuel primer valve to the open position and opens the

motor-driven crossfeed separation valve. This allows fuel to flow through the manifold into the No. 2 fuel tank to

remove any trapped air. Releasing the button actuates the primer valve to the closed position and closes the crossfeed

separation valve. Power for the system comes through the ENGINE CROSSFEED VALVES PRIME circuit breaker

on the copilot side circuit breaker panel.

2.6.7 External Tanks

Two all-metal external fuel tanks are mounted under thewings on pylons between the inboard and outboard engines.

The tanks are partially compartmented for center-of-gravity control. All fuel flows into the center compartment

through check valves. A surge box in the tank center compartment contains a forward and an aft boost pump,

providing dual reliability and an increased fuel dumping rate if both pumps are operated during fuel dumping. Both

pumpshaveoverridingoutputpressures that,undernormaloperation, ensuredepletion offuel fromtheexternaltanks

before the main tanks are affected.

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ORIGINAL

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2.6.8 Fuel Dump System

A fuel dump system is provided to enable all fuel, except approximately 1,600 pounds each from the No. 1 and 4

wing tanks, 1,500 pounds each from the No. 2 and 3 wing tanks, and 65 pounds from each external tank, to be dumped

overboard.Eighttwo-position(OFF,DUMP)toggleswitchesforthewingtanksarelocatedon theenginefuelcontrol

panel. Two-position (NORM, DUMP) guarded toggle switches and two rotary interconnect switches are located on

theauxiliaryfuelcontrolpanel(Figure2-24).Thedumprateisapproximately3,900 poundsperminutefromthewing

tanks with all pumps operating. All tanks feed into a common manifold in the wing, then to a dump mast in each

wingtip. Check valves at each tank dump outlet prevent reverse flow. The four main wing tanks have individual

integral pumps specifically for fuel dumping. The two auxiliary and two external tanks use the same pumps for

dumping that are used for normal boost operation. Actuation of the tank dump switch will open the tank dump valve

and turn the pump on for the selected tank (see FO-1). To complete the dump operation, the interconnect switches

on the auxiliary fuel panel are turned to the flow condition and the two guarded dump switches on the auxiliary fuel

control panel are placed in the DUMP position. When the auxiliary tank dump switches are actuated, the auxiliary

tank crossfeed valves will close. Refer to Chapter 11 for operation of the wing tank fuel dump system and to Part

VIII for operation of the fuselage tank dump system. All pumps are powered by three-phase, 200/115-volt, 400-Hz

ac. Power for the motor-operated dump valves is supplied from the essential dc bus through the LH and RH FUEL

DUMP VALVES circuit breakers on the copilot side circuit breaker panel. The dump pumps in the four main wing

tanks are supplied power from the main ac bus through the DUMP PUMPS circuit breakers on the copilot upper

circuit breaker panel.

Figure 2-23. Fuel Control Panel

ORIGINAL

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01-75GAL-1

2.6.9 Fuel Strainer and Heater Unit

A combination fuel strainer and heater is located in the right side of each nacelle. Heat is transferred from engine oil

to the fuel in the heater unit, and the temperature is controlled thermostatically.

2.6.10 Fuel System Controls

Controls for normal in-flight management of engine fuel are located on the fuel control panel (see Figure 2-23).

2.6.10.1 BOOST PUMP Switches

Ten BOOST PUMP switches are located on the fuel control panel. The No. 1, 2, 3, and 4 fuel tank boost switches

control the internal boost pumps for their respective tanks. The left and right AUX TANK PUMP switches control

the pump in each of the auxiliary tanks. Four BOOST PUMP switches, one forward and one aft for each external tank,

control the pumps in the external tanks. All of the boost pumps are powered by three-phase, 115/200volt, 400-Hz

ac. The No. 1 tank pump is supplied power through the FUEL BOOST PUMP TANK NO. 1 circuit breakers on the

pilot upper circuit breaker panel left-hand ac bus. The No. 2 tank pump is supplied power through the FUEL BOOST

PUMP TANK NO. 2 circuit breakers on the pilot side circuit breaker panel essential ac bus. The No. 3 tank pump

is supplied power through the FUEL BOOST PUMP TANK NO. 3 circuit breakers on the copilot upper circuit

breaker panel main ac bus. The No. 4 tank pump is supplied power through the FUEL BOOST PUMP TANK NO.

4 circuit breakers on the pilot upper circuit breaker panel right-hand ac bus. The left- and right-hand auxiliary fuel

tank boost pumps are supplied power through the AUXILIARY TANK LH and RH circuit breakers on the copilot

uppercircuit breakerpanel main acbus. Theright- and left-hand external fuel tank forward boost pumps aresupplied

power through the RH and LH EXT PUMP (FWD) circuit breakers on the pilot upper circuit breaker panel from the

RH and LH ac buses respectively. The right- and left-hand external fuel tank aft boost pumps are supplied main ac

power through the RH and LH EXT TANK PUMP (AFT) circuit breakers on the copilot upper circuit breaker

panel.

2.6.10.2 CROSSFEED VALVE Switches

Eight CROSSFEED VALVE switches are located on the fuel control panel (see Figure 2-23). These two-position

rotary switches route 28-Vdc through the ENGINE CROSSFEED VALVES circuit breakers on the copilot side

circuit breaker panel to the motor-operated crossfeed valves. When the switches are placed in the flow position

(switch markings aligned with the fuel control panel markings), the valve motors are energized to open the valves.

When the switches are placed in the no-flow position (switch markings at right angles to the panel markings), the

valve motors are energized to close the valves. In case of power failure, the valves hold the last energized position.

2.6.10.3 Bypass Valve Switches

Two bypass valve switches are located on the fuel control panel to permit an alternate path for fuel from the left and

right auxiliary and external fuel tanks if crossfeed valves fail to open. These two-position rotary switches route power

to motor-operated bypass valves.When theswitches areplaced inbypass position(switch markingsaligned withfuel

control panel markings), valve motors are energized to open the valves and allow external tank fuel to be crossfed

or jettisoned through the auxiliary tank crossfeed or jettison valves, and vice versa. The bypass valves may be used

to jettison main tank fuel in the event of main tank dump valve/pump failure. When switches are placed in the off

position (switch markings at right angles to panel markings), valve motors are energized to close the valves. In case

of power failure, the valves hold the last energized position. The bypass valve switch receive power from the essential

dc bus through the FUEL MANAGEMENT ENGINE CROSSFEED VALVES 1 & 4 (LEFT VALVE) and 2 & 3

(RIGHT VALVE) circuit breakers located on the pilot side circuit breaker panel.

2.6.10.4 Crossfeed Separation Valve Switch

The CROSSFEED separation switch is located on the fuel control panel (see Figure 2-23). The crossfeed separation

valve is provided in the crossfeed manifold system to permit additional control on fuel routing. With the crossfeed

separation valve closed, the left wing tanks supply fuel to engine Nos. 1 and 2, while the right wing tanks supply fuel

toengineNos.3and4.Thisprocedureensuresamoreeven fuelconsumption whenoperating fromtheauxiliarytanks

through the crossfeed manifold. Since there may be aslight variation in boost pump pressureand ifboth pumps were

supplying the manifold, the pump operating at the highest pressure would feed the manifold if not prevented by the

separation valve. However, the crossfeed separation valve is used primarily for fuel balancing between the left- and

right-hand fuel tanks and when crossfeeding from the fuselage tank. The crossfeed separation valve is electrically

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actuated by 28-Vdc power received from the essential dc bus through the ENGINE CROSSFEED VALVES PRIME

circuit breaker, located on the copilot side circuit breaker panel.

2.6.10.5 Fire Emergency Control Handles

Five fire emergency control handles, one for each engine and one for the auxiliary power unit, are mounted on the

fire emergency control panel (see Figure 2-17). These control handles route 28-Vdc power to the motor-operated,

engine firewall fuel shutoff valves and to the motor-operated, auxiliary power unit fuel supply shutoff valve. In case

of power failure, valves hold the last energized position. Circuit protection is provided by the FIRE SHUTOFF

VALVES circuit breakers on the copilot side circuit breaker panel and the APU CONTROL circuit breaker on the

pilot side circuit breaker panel. Other functions of the handles are described under paragraph 2.3.1.2.

2.6.11 Fuel System Indicators

Quantity gauges and warning lights are located on the fuel control panel (see Figure 2-23) and on the auxiliary fuel

control panel (see Figure 2-24) to give the crew a continuous, visual indication of the status of the fuel system. For

additional information on fuel indicators, refer to paragraph 2.1.8.

2.6.11.1 TOTAL FUEL QUANTITY Indicator

A TOTAL FUEL QUANTITY indicator is located in the center of the fuel control panel (see Figure 2-23). The

indicator continuously shows the total fuel quantity (in pounds) in the fuel tanks when the single-point refueling

master switch is in the OFF position. When the master switch is in any position other than OFF, the TOTAL FUEL

QUANTITY indicator is deenergized. The indicator is tested when an individual tank quantity indicator is tested as

described in paragraph 2.6.11.2. The TOTAL FUEL QUANTITY indicator receives single-phase, 115-volt, ac power

from the ac instruments and fuel control bus through the FUEL QUANTITY TOTALIZER circuit breaker on the pilot

lower circuit breaker panel.

2.6.11.2 FUEL QUANTITY Indicators and Test Switches

Note

On aircraft 165313 and up there is no analog bar graph on the flight

station indicators.

FUEL QUANTITY indicators are located on the fuel control panel (see Figure 2-23). Each tank indicator provides

a continuous visual display of the pounds of fuel contained in that tank. A liquid crystal display shows fuel quantity

in both digital and analog format. A three-digit numeric readout shows fuel quantity in thousands of pounds. An

analog bar display around the periphery of the dial shows the percentage of total fuel capacity of that tank. Each

illuminated small mark represents 2 percent of tank capacity. The major marks for each 10 percent of tank capacity

are always illuminated. Separate conductors in the aircraft wiring harness provide a binary code that tells the

individualindicatorcomputertowhichtankitisconnected.Theindicatoris thenproperly scaledforthattank. Aletter

(M, E, A) and a number (1, 2, 3, 4) appear at the bottom of the indicator display that identifies the tank. The main

tanks are identified as M1, M2, M3, and M4; the auxiliary tanks are identified as Al and A2, and the external tanks

are identified as El and E2. All eight indicators on the fuel control panel are interchangeable. Each indicator has

computerized built-in-test equipment. The BITE is designed to test the entire fuel quantity indicating system, not just

the indicator. The BITE functions are continually processed during normal system operation. Additional test

functions are performed at power-up and when the IND TEST pushbutton is depressed. The results of the BITE test

are displayed in numerical code by the indicator. Pressing the IND TEST pushbutton on the fuel control panel

adjacent to the indicator causes the display reading to decrease toward zero. Releasing the IND TEST pushbutton

causes all segments of the digital display, analog display, tank identification and error code to illuminate for

approximately 4 seconds, after which all segments will extinguish for approximately 4 seconds. Then any error code

that has been detected since power-up will be displayed for approximately 4 seconds until all errors found by thebite

test are displayed. During normal operation, all fault conditions are detected within 10 seconds after occurrence, and

the error code is displayed on the indicator. If thedetected fault condition should return to normal, theerror codewill

be removed and the normal fuel quantity display will automatically resume. The following error codes may be

displayed:

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ERROR CODE

DESCRIPTION

FUEL QUANTITY INDICATION

Open or missing tank unit

Zero

Tank unit leakage

Displayed at reduced accuracy

Tank unit Hi-Z fault

Zero

Compensator Lo-Z fault

Displayed at reduced accuracy

Tank unit Lo-Z fault

Zero

Compensator fault

Normal display at reduced accuracy

Shorted tank unit

Zero

Compensator leakage or out of

Displayed at reduced accuracy

range

Indicator internal failure

Zero or blank

Internal calibration failure

Zero

Blank Display

Indicator failure

Blank

The BITE for the TOTAL FUEL QUANTITY indicator is continually processed during normal operation of the

system. The indicator displays error codes that are either detected within the indicator itself or received from any of

the eight fuel quantity indicators. The TOTAL FUEL QUANTITY indicator is tested when the IND TEST

pushbutton for any individual fuel quantity indicator is depressed. When an IND TEST pushbutton is depressed, the

TOTAL FUEL QUANTITY indicator reading will decrease in response to the decrease in the individual tank fuel

quantity indicator. The fuel quantity indication will continue to decrease to zero until the IND TEST pushbutton is

released. When the pushbutton is released, all segments of the total fuel quantity indicator display will illuminate

for approximately 4 seconds. Then all segments will be extinguished for 4 seconds. When the TOTAL FUEL

QUANTITY indicator is in the BITE mode, it will not display any type of error code. If an individual tank fuel

quantity indicator has more than one error code to display during the BITE test, the total fuel quantity indicator will

return to normal fuel quantity display. The error codes are continuously displayed whenever the fault is present. If

more than one fault is present, each fault will be sequentially displayed for 1 second.

Note

Tank indicator error codes indicate an error in the individual fuel tank

quantity indicator. A communications errorcodeis displayed when nodata

or invalid data is received from a tank indicator. The number appearing

after the first letter identifies the tank. For example, M2E0 indicates No.

2 main tank communications error.

The following error codes are detected by the TOTAL FUEL QUANTITY indicator and the repeater indicators.

ERROR

CODE

DESCRIPTION

M—E

Main tank indicator error

A—E

Auxiliary tank indicator error

E—E

External tank indicator error

M — E0

Main tank indicator communication error

A — E0

Auxiliary tank indicator communication error

E — E0

External tank indicator communication error

Blank Display

Total fuel quantity indicator failure

Display is blank with error code; indicator internal failure

Display shows zero with error code; calibration failure

The fuel quantity indicating system receives 115-Vac power from the ac instrument and engine fuel control bus.

Circuit protection is provided by the FUEL QUANTITY circuit breakers on the pilot lower circuit breaker panel.

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2.6.11.3 Auxiliary Fuel Tank Magnetic Sight Gauge

An auxiliary fuel tank magnetic sight gauge is located on the underside of the wing center section for each auxiliary

fuel tank. The magnetic sight gauge consists of three components: a mounting base and outer tube, a float, and a gauge

stick. The mounting base is attached to the lower surface of the auxiliary fuel tank with the outer tube secured to the

mounting base. The float rides on the outside of the tube and has magnets in its inner diameter. The gauge stick is

contained within the outer tube, has magnets on its upper end, markings to indicate fuel quantity, and latches on the

lower end that latch into the mounting base. The gauge stick markings indicate fuel quantity and are marked from

5 to 59 in 500-pound increments.

2.6.11.4 AUX and EXT TANK EMPTY Lights

Two AUX TANK EMPTY lights and two EXT TANK EMPTY lights are located on the fuel control panel (see Figure

2-23), in the flight station. If the BOOST PUMP switch associated with a given auxiliary or external tank is positioned

to ON and there is no source of higher pressure to that side of the manifold, the associated tank empty light will be

illuminated whenever output flow pressure is below approximately 23 psi. Illumination of the light indicates either

depleted tank quantity or an inoperative boost pump or (in the case of the external tanks only) failure of the fuel level

control valve in the open position. The tank empty lights receive 28-Vdc power from the essential dc bus through

the respective FUEL DUMP VALVES AUX or EXT circuit breakers on the copilot side circuit breaker panel. Refer

to Part VIII for discussion of the fuselage tank empty light.

2.6.11.5 REFUEL PANEL ON Light

AREFUEL PANEL ON light is located on thefuel control panel (seeFigure 2-23). This press-to-test light will come

on anytime the single-point refueling MASTER switch is in any position other than OFF. Power for the REFUEL

PANEL ON light is supplied from the essential dc bus through the WARNING LIGHTS TEST circuit breaker on

the copilot lower circuit breaker panel.

2.6.11.6 Fuel STRAINER BYPASS OPEN Light

A fuel STRAINER BYPASS OPEN light is located on the fuel control panel (see Figure 2-23). A pressure switch

on the strainer unit actuates the light when the pressure drop across the strainer exceeds approximately 4 PSID. Power

for the bypass light is supplied from the essential dc bus through the ENG FUEL STRAINER BYPASS INDICATOR

circuit breaker on the copilot side circuit breaker panel.

2.6.11.7 Fuel LOW PRESS Warning Lights

Four fuel LOW PRESS warning lights are located on the fuel control panel (see Figure 2-23). Each light goes on when

pressure in the fuel supply line to the engine pump drops below approximately 8.5 psi. When a light goes on, it

indicates a possible booster pump failure, valve failure, fuel line failure, or a malfunctioning pressure switch. The

lights are energized by 28-Vdc power from the essential dc bus through the LOW PRESSURE LIGHTS circuit

breaker on the copilot side circuit breaker panel.

2.6.11.8 MANF PRESS Indicator

AMANFPRESSindicator,locatedonthefuelcontrolpanel(seeFigure2-23),indicatesfuel pressurein thecrossfeed

manifold. The indicator is used to check fuel boost pumps before starting engines and used during flight only to

determine if a pump is operating. This indicator is electrically connected to a fuel pressure transmitter. The transmitter

measures the pressures of the crossfeed manifold. Thus, when the fuel boost pumps are turned on individually, the

pressure supplied the crossfeed system by any pump is measured by the transmitter and shown by the indicator.

Single-phase, 26-Vac to operate the pressure indication system is supplied by the No. 1 instrument transformer.

Circuit protection is provided by the FUEL PRESSURE INDICATOR fuse on the pilot lower circuit breaker panel.

2.6.12 Single-Point Refueling and Defueling System

A single-point refueling and defueling system enables all normal refueling and defueling operations to be

accomplished through a single receptacle located in the aft end of the right wheelwell fairing. Controls and indicators

for all tanks except the fuselage tank are installed on the single-point refueling control panel located immediately

above the receptacle (see Figure 2-25). Controls and indicators for the fuselage tank are mounted on the auxiliary

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01-75GAL-1

fuel control panel (see Figure 2-24). When refueling, fuel enters the tanks by way of the refueling manifold. A dual

float valve in each tank shuts off the flow when the tank is filled to its single-point refueling capacity. Defueling can

be accomplished by using the tank boost pumps or the dump pumps. When using the boost pumps, the fuel flow is

through the crossfeed manifold, through the ground transfer valve to the refueling manifold, and out the refueling

receptacles. Defueling flow, when using the dump pumps, is through the dump line to the refueling manifold and

out the refueling receptacles. Fuel can be transferred between fuel tanks when the aircraft is on the ground by using

the tank boost pumps or dump pumps and single-point refueling controls. A surge suppressor is located in the

refueling line to absorb pressure surges during offloading. A surge suppressor pressure gauge is located on the surge

suppressor and may be observed through the starboard air deflector door.

2.6.12.1 Single-Point Refueling and Defueling System Controls and Indicators

2.6.12.1.1 MASTER Switch

Amasterswitchforthesingle-point refuelingsystem islocated onthesingle-pointrefueling controlpanel (seeFigure

2-25). The switch is a six-position (DRAIN, DEFUEL, OFF, PRE-CHKSEC, REFUEL & GRD TRANS, PRE-CHK

PRIM) rotary-type by which the system function is selected. Placing the MASTER switch in the REFUEL & GRD

TRANS position supplies power to the tank selector switches and the ground transfer switch, permitting selective

(OPEN, CLOSE) operation of the tank fill valves and the ground transfer valve. Placing the MASTER switch in the

DEFUEL position supplies power to operate the ground transfer valve only. The tank fill valves cannot be opened

when the MASTER switch is in the DEFUEL position. Placing the switch in either the PRE-CHK PRIM position

or the PRE-CHK SEC position interrupts power to the corresponding solenoid in the tank fill valves, closing the fill

valves and simulating a tank-full condition, thus providing a check on the automatic operation of the tank fill valves.

In both the PRE-CHK PRIM and the PRE-CHK SEC positions, power is supplied to the ground transfer switch,

permitting operation of the ground transfer valve. In the DRAIN position, power is supplied to open the drain valve

and to operate the drain pump. Power is also supplied directly to the offload valve, bypassing the OFFLOAD VALVE

switch(closingthevalve),andpowerisremovedfromthetankselectorswitches(renderingtheswitchesinoperative).

In all positions except OFF, the single-point refueling panel fuel quantity gauges are energized. The refueling system

receives 28 Vdc from the main dc bus through the REFUELING PANEL circuit breakers on the copilot lower circuit

breaker panel.

2.6.12.1.2 Tank Selector Switches

Each of the selector switches, one for each tank, located on the single-point refueling control panel (see Figure 2-25),

is a two-position (OPEN, CLOSE), rotary-type switch through which power is supplied to the solenoids of the

associated tank fill valve. The switches can operate the valves only while the fueling control MASTER switch is in

REFUEL & GRD TRANS.

2.6.12.1.3 GROUND TRANSFER Switch

The GROUND TRANSFER switch, located on the single-point refueling control panel (see Figure 2-25), is a

two-position (OPEN, CLOSE), rotary-type switch, used to control the ground transfer valve. The MASTER switch

mustbeintheDEFUEL,PRE-CHKSEC,REFUEL&GRDTRANS,orPRE-CHKPRIMpositionbeforetheground

transfer valve will operate. When the master switch is in the OFF or DRAIN position, the GROUND TRANSFER

switch is bypassed, and the valve is energized to the closed position.

2.6.12.1.4 OFF-LOAD VALVE Switch

The OFF-LOAD VALVE switch, located on the single-point refueling control panel (see Figure 2-25), is a

two-position (CLOSE, OPEN), toggle-type switch used to control the offload valve when the MASTER switch is

inanypositionexceptOFForDRAIN.WhentheMASTERswitchis intheOFForDRAINposition, theOFF-LOAD

VALVE SWITCH is bypassed and the offload valve is energized to the closed position. The OFF-LOAD VALVE

switch must be in the OPEN position to refuel and defuel.

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01-75GAL-1

2.6.12.1.5 FUEL QUANTITY Indicators

One FUEL QUANTITY indicator for each fuel tank is located on the single-point refueling control panel (see Figure

2-25). All the indicators register tank fuel quantity in pounds and are energized when the MASTER switch is at any

setting other than OFF. The fuel quantity indicators have a liquid crystal display that shows fuel quantity in pounds

on a digital display and the percentage of total fuel capacity on an analog display. The single-point refueling control

panel indicators are similar to those installed on the flight station fuel control panel except for a zero adjustment screw

located on the aft end of the flight station indicators. The single-point refueling control panel repeater indicators are

tested by depressing the corresponding flight station fuel control panel IND TEST pushbutton with the single point

refueling system MASTER switch in any position other than OFF. An E0 error code on the repeater indicator

indicates that no signal is being supplied from the primary indicator in the flight station. An E-error code on the

repeater indicator indicates that the primary indicator is presently displaying an error code. The repeater indicator

error codes are the same as the total fuel quantity indicator error codes. With an IND TEST pushbutton depressed,

both primary and repeater indicator numeric readouts count down to zero. When the IND TEST pushbutton is

released, the LCD segments illuminate for 4 seconds, then go blank for 4 seconds. The repeater indicator then

displays fuel quantity. The indicators receive single-phase, 115-Vac power from the ac instruments and fuel control

bus through the FUEL QUANTITY TOTALIZER circuit breaker on the pilot lower circuit breaker panel.

2.6.12.1.6 INTER CONN Switches

Two rotary, two-position (flow, no-flow) interconnect valve switches are provided on the auxiliary fuel control panel.

The switches enable dumping to be accomplished from the wing tanks and external tanks. When the switch is in the

FLOW position, fuel is supplied from the wing, external and fuselage tanks; however, the fuel supplied from the

fuselage tank must be by the FUS TANK CROSSFEED switch. Power for operation of the interconnect valves is

supplied from the essential dc bus through the INTERCONNECT VALVE RH and LH circuit breakers on the copilot

side circuit breaker panel. Power for operation of the interconnect valve relay is supplied from the essential dc bus

through the INTERCONNECT VALVE CONTROL circuit breaker on the copilot side circuit breaker panel. The

interconnect valves must be open for dumping fuel from the wing and external tanks.

2.6.12.1.7 Fuel Dump Switches

Two two-position (NORM, DUMP) fuel dump switches are provided on the auxiliary fuel control panel for dumping

fuel from the wing tanks. When the switches are placed in DUMP (unguarded) position, the dump valves will open.

Wing tank fuel can be dumped by placing the desired wing tank dump switch to the DUMP position and turning the

interconnect switches to the FLOW position. Power to the dump switches is supplied from the essential dc bus

through the FUEL DUMP VALVES LH and RH circuit breakers on the copilot side circuit breaker panel.

2.6.12.2 Refueling Procedure

Refer to Chapter 3.

2.7

ELECTRICAL POWER SUPPLY SYSTEMS

Note

Never use circuit breakers as switches. Circuit breakers should be pulled

only during emergencies or maintenance. (See Figures 2-34 through 2-41

for circuit breaker panels.)

All internal electrical power for aircraft use comes from five ac generators (see Figure 2-26), or from the battery. Each

engine drives one 40-Kva, ac generator, and a 40-Kva generator is driven by the APU. The APU-driven generator

is the same type as the engine-driven generator. Power from these ac generators is used to provide electrical power

for aircraft use: 28-Vdc; 200/115-volt, 400-Hz, 3-phase primary ac; and 115-volt, 400-Hz, single-phase, secondary

and primary ac. The four engine-driven ac generators are connected through transfer contactors (relays) to four ac

ORIGINAL

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01-75GAL-1

buses: the left-hand ac bus, the essential ac bus, the main ac bus, and the right-hand ac bus. The transfer system

operates in such a manner that any combination of two or moreof theengine-driven acgenerators will power all four

of the buses. On aircraft 165313 and up, two bus switching units conduct ac power from the unregulated ac buses

to two additional (essential and main)avionics acbuses. (SeeFigure2-29foradistribution schematicofthesebuses.)

The bus switching units provide uninterrupted ac power to critical avionics components through the avionics buses

by monitoring the various sources and switching rapidly between thesesources as required. TheBSUs arecontrolled

by BSU 1 and BSU 2, two-position (OFF, ON) toggle switches on the overhead electrical control panel. (See Figure

2-28.) If only one generator is operating, it will power only the essential ac bus and the main ac bus. Placing the APU

generator control switch in the APU position energizes the APU generator contactor, which connects the APU

generator to the essential ac bus. The APU-driven ac generator normally will power only the essential ac bus. The

APU generator will also power the main ac bus provided there are no engine-driven generators supplying power and

the AC BUS TIE switch is placed to the ON position. This feature is for ground use only. The APU generator, voltage

regulator, and generator control are the same as the engine-driven generator, voltage regulator, and generator control.

(Refer to paragraph 2.19.) Combinations of operating generators and the buses that they power are shown on the ac

bus power sources chart (see Figure 2-32). All in-flight controls for operation of the electrical system are located on

the overhead electrical control panel (see Figure 2-28). Circuit breakers are shown in Figures 2-34 through 2-41.

2.7.1 Bus Switching Units (Aircraft 165313 and Up)

Two bus switching units (BSU 1 and BSU 2) provide continuous no-break power to the essential and main avionics

ac buses for critical avionics systems. BSU No. 1 supplies the essential avionics ac bus, and BSU No. 2 supplies the

main avionics ac bus. During normal operation with generator power, BSU No. 1 and No. 2 both receive power from

the essential and main ac buses. For BSU No. 1, the essential ac bus is the primary source and the main ac bus is the

alternate source. For BSU No. 2, the main ac bus is the primary source and the essential ac bus the alternate source.

When external ac power is being used, BSU No. 1 receives power from the left-hand ac bus only and BSU No. 2

receives power from the right-hand ac bus only. When the BSU switches are OFF or the BSUs are in the bypass mode,

the essential and main avionics buses are powered directly from their respective primary source. Each BSU receives

logic inputs that provideelectrical system status. It monitors theengine generatorline contactors, the APU generator

contactor, and the ac external power contactor, thus receiving advance information of impending bus transfers. It also

monitors voltage, frequency and current, and selects the best source of power for output to the respective avionics

bus.Duringnormalstableoperations,BSUNo.1selectspowerfromtheessentialacbusandBSUNo.2 selectspower

from the main ac bus.

Note

With the APU generator connected to the essential ac bus, programmed

logic in BSU No.1 does not recognize the APU on status. Simultaneous

disconnecting or reconnecting engine generator Nos. 2 and 3 when in this

configuration will cause momentary interruption of power to the essential

ac avionics bus.

When ac power is first applied to the aircraft with the BSU switches OFF, the BSUs are in the bypass mode with the

BSU BYPASS lights illuminated. When theBSU switches are placed to ON, the BSUs each perform a 5-secondBIT

check. After a successful BIT check, the BSU BYPASS lights will extinguish and the BSUs will connect the avionics

busestotheprimaryoralternatesource, asavailable, ifon generatorpower, orto theleft-hand andright-hand acbuses

if external ac power is selected. With external ac power selected, when the first generator is placed on line, external

ac power is removed from the essential and main ac buses as their load is assumed by the generator. External power

is then removed from the left- and right-hand ac buses after a delay of 0.5 seconds, and the BSUs then transfer to the

primary or alternate sources, as available. As additional generators are brought on line, if the BSU detects a power

interruption or a condition that would cause a degradation in output power quality, it will transfer the output to the

alternate source until the primary source stabilizes. After a 7-second delay, the BSU will switch back to the primary

source, provided it is stable.

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01-75GAL-1

Note

In order to ensure a smooth transition of power to the main avionics ac bus,

place the AC BUS TIE switch to ON prior to switching from external ac

power to APU generator power.

The BSUs receive 28-Vdc control power through BSU 1 and BSU 2 two-position (OFF/ON) toggle switches on the

overhead electrical control panel. When placed in the OFF position, the BSU reverts to a bypass condition that

supplies power to the respective avionics bus from its primary ac bus. During normal operation, the BSU may

automatically go into the bypass mode if it fails to pass BIT. Two advisory lights illuminate to indicate the BSUs

are in the bypass mode either because the BSU has failed BIT or the BSU switch has been placed to OFF. In the event

of an overload condition on an avionics bus, the affected BSU will go into bypass mode and the respective BSU 1

or 2 BYPASS light will illuminate if the overload is not cleared in a specified time. The specific time varies from

5 minutes at 150-percent BSU overload, to 5 seconds at 200-percent overload, to 2 Hz for a short to ground. The delay

allows time for circuit protection devices to trip or open to relieve the overload. These switches can also be used at

any time to try to reset the BSU if it should automatically go into the BYPASS or OFF mode. It takes approximately

10 seconds for the BSU to reset. The BSU will remain in the bypass mode and will not reset until all power to the

BSU has been removed and reapplied.

2.7.2 External Power Provisions

Note

The 200/115-volt, three-phase, 400-Hz ac external source should have a

capacity of 40 Kva; its phase rotation must be A-B-C. The 28-Vdc external

source should have a capacity of 400 amperes.

Both dc and ac external power receptacles are located on the left side of the fuselage just aft of the battery

compartment. Dc power from the external source is supplied through two current limiters to the main dc bus. Any

dc electrically operated equipment on the aircraft, except equipment connected to the battery bus, can be supplied

from an external dc power source. The battery is disconnected from all dc buses except the battery bus when external

dc power is being used. When an external ac power source is connected to the aircraft, power is supplied to all ac

buses, to the dc buses through transformer-rectifier units, and to the battery bus to charge the battery if the DC

POWER switch is in the BATTERY position.

Note

The APU generator switch must be in the OFF position before external ac

power can be fed into the aircraft system.

2.7.3 Primary Ac System

Power for the primary ac system is supplied by five ac generators (see Figure 2-26). This power supply is also used

to operate the secondary ac systems and the dc system.

2.7.4 Primary Ac System Controls

The ac system controls, with the exception of a manual reset lever on each generator control panel, are located aft

of the overhead electrical control panel (see Figure 2-28) in the flight station. The generator control panels are located

in racks under the flight station and are accessible from the cargo compartment.

2.7.4.1 Generator Switches

The generator switches consist of five four-position rotary-type switches. When aswitch is in theON position (knob

stripe aligned with panel stripe), a relay closes contacts to connect the generator to the buses if the generator is

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operating normally. The distribution of generated power to the various buses under all conditions of generator

operation is shown in Figures 2-27 and 2-33. All engine generators are off the line when the low-speed ground idle

buttons are engaged; therefore, the APU generator must be used to provide power to the essential ac bus, which is

the only bus it supplies. The APU generator may also supply the main ac bus, provided there are no engine-driven

generators supplying power and the ac BUS TIE switch is placed in the ON position. This feature is for ground use

only. When a generator switch is placed in the OFF or (for aircraft 165313 and up) OFF/RESET position, the relay

disconnects the generator from the system. If the switch is turned to TRIP, the field circuit of the generator is opened

by a field relay to remove generator excitation. No voltage is then produced by the generator. The RESET position

of the switch is used to operate the field relay to its reset position after it has been tripped. The relay then closes the

generator field circuit to allow the generator to build up voltage. On aircraft prior to 165313, the RESET position

of the generator switch knob is spring loaded. The generator switch knob must be pulled out to move it to the TRIP

position.

2.7.4.2 Generator Disconnect Switches

Each engine-driven generator is provided with a spring-loaded, two-position (OFF, DISC) guarded switch. When

the switch is held in the DISC position (approximately 2 seconds), a direct short in the firing mechanism causes the

fused portion of the plunger to burn through and be actuated by spring tension. As the plunger of the firing mechanism

passes over the generator disconnect fired light switch, the generator DISC FIRED light will illuminate, indicating

the firing mechanism has been fired. The plunger then engages a wing on the generator stub shaft causing it to shear.

The generator cannot be reconnected in flight since a new stub shaft must be installed. Power to the switches is

supplied from the essential dc bus through the four GENERATOR DISCONNECT PWR circuit breakers on the pilot

upper circuit breaker panel.

2.7.4.3 Generator Disconnect Test Switch

A spring-loaded, two-position OFF, TEST GEN DISC switch is provided to check the continuity of the firing

mechanism. If the continuity check is good, the generator DISC FIRED lights will illuminate. Power to the switch

is supplied from the essential dc bus through the GENERATOR DISCONNECT PWR circuit breaker on the pilot

upper circuit breaker panel.

2.7.4.4 Ac BUS TIE Switch

Atwo-position(OFF,ON)acBUSTIEswitchontheoverheadelectricalcontrolpanelprovides ameans forpowering

the main ac bus from the APU generator during ground operation with no engine-driven generators supplying power.

The switch receives 28-Vdc power from the essential dc bus through the APU GEN ESS TO MAIN 53K circuit

breaker in the main ac distribution panel.

2.7.4.5 Ac External Power Switch

A two-position, ac external power switch is located immediately below the LH ac bus loadmeter on the overhead

electrical control panel. The OFF position of the switch disconnects external power from the ac distribution system.

The external power position (stripe on knob aligned with stripe on panel) connects external power to the ac

distribution system. The ac external power switch receives 28-Vdc control power from the ac external power source

throughtheaircraftexternalacpowerunit,or24-Vdccontrolpowerdirectlyfromtheacexternal powercart dcbattery

through the aircraft ac EXT PWR CONTROL circuit breaker in the aircraft battery compartment.

Note

Anoverridesolenoidinthesystemispoweredfromthebatteryandwillturn

the switch off if the APU generator control switch is on, if the ac power is

not in the correct phase sequence, if any engine generator is on the line, or

if the external power plug is not in the receptacle.

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2.7.4.6 VOLTAGE & FREQUENCY SELECTOR Switch

Aseven-position (ENG GEN NO. 1, ENG GEN NO. 2, ENG GEN NO. 3, ENG GEN NO. 4, APU GEN, EXT PWR,

COPLT INV φA AC INST & ENG FUEL CONT INV φC) rotary switch is used to isolate a chosen source of ac power

for measurement.

2.7.4.7 PHASE SELECTOR Switch

A three-position PHASE SELECTOR switch selects one of the three phases of ac generator output to be measured

by the ac meters on the panel. Placing the switch in a given position determines which phase is measured by the five

ac loadmeters, the ac voltmeter, and the frequency meter on the overhead electrical control panel. The switch can be

used in conjunction with the VOLTAGE & FREQUENCY SELECTOR switch to test a given ac power source.

2.7.4.8 Primary Ac System Indicators

Indicators for the primary ac power system are located in the overhead electrical control panel (see Figure 2-28) in

the flight station.

2.7.4.9 Generator-Out Indicator Lights

Each generator is provided with a generator-out press-to-test indicator light. This light will illuminate when the

generator control switch is in the ON position and one or more of the following conditions exist: the generator is not

developing sufficient voltage, the generator output is below proper frequency, or the field-trip relay has opened the

field circuit of the generator. The field relay will trip when the generator switch is turned to the TRIP position, when

the generator output voltage is too high, or when a fault exists in the generator output circuit.

2.7.4.10 Ac Loadmeters

Five ac loadmeters, one for each generator, give a continuous indication of the percent of rated current flow of any

one PHASE SELECTOR switch from their respective generators.

2.7.4.11 AC BUS OFF Indicators

The AC BUS OFF warning system consists of four red press-to-test warning lights and four indicator relays. The

warning lights are located on the flight station overhead electrical control panel and the relays are located behind the

LH and RH distribution panels. The lights are labeled ESS AC OFF, MAIN AC OFF, LH AC OFF, and RH AC OFF.

On aircraft 165313 and up there are six red press-to-test warning lights and six indicator relays. Labeled ESS AC

OFF, MAIN AC OFF, LH AC OFF, RH AC OFF, ESS AV AC OFF, and MAIN AV AC OFF. The indicator relays

receive power from their respective ac bus. The warning lights receive 28-Vdc power from the isolated dc bus through

the AC BUS OFF IND circuit breaker on the pilot side circuit breaker panel.

2.7.4.12 AC VOLTMETER

The AC VOLTMETER can be used to measure the output voltage of the generator or inverter that has been selected

with the VOLTAGE & FREQUENCY SELECTOR switch. Each of the three phases of generator output, or the

appropriate phase of inverter output, can be measured by selectively positioning the PHASE SELECTOR switch.

2.7.4.13 Frequency Meter

A frequency meter permits measuring the frequency of the output power of the generator selected with the VOLTAGE

& FREQUENCY SELECTOR switch. Each of the three phases of the generator output power can be measured by

selecting the appropriate position on the PHASE SELECTOR switch.

2.7.4.14 Ac External Power-On Indicator Light

An ac external power-on (EXT AC PWR) press-to-test indicator light is energized by dc power through small pins

in the ac external power receptacle and through the closed contacts of a phase sequence relay on the lower main ac

distribution panel when the relay is energized. The phase sequence relay is energized when three-phase external ac

power with correct phase sequence and no open phases is connected to the aircraft.

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2.7.4.15 Generator Disconnect Fired Indicator Lights

Each engine-driven generator is provided with a generator disconnect fired (DISC FIRED) indicator light that will

illuminate when one of the following conditions exists: a generator disconnect switch is held in the DISC position

and the firing mechanism is fired, or when the generator disconnect test switch is held in the TEST GEN DISC

position. Power to the lights is supplied from the essential dc bus through the GENERATOR DISCONNECT circuit

breaker on the pilot upper circuit breaker panel.

2.7.4.16 Generator Bearing-Failure Indicator Lights

Each engine-driven generator is provided with a generator bearing failure (FAILED BRG) warning indicator light.

Each generator stator contains a soft wire winding that grounds out the stator when contact is made with the rotor.

When this occurs, a circuit is completed for illumination of the generator bearing-failure indicator light. Once the

indicator light is illuminated, it will remain illuminated until dc power to the circuit is removed. The power for the

lights and holding circuit is supplied from the essential dc bus through two GEN BRG FAIL circuit breakers on the

pilot upper circuit breaker panel.

2.7.4.17 Ac BUS TIE Switch On Light

An amber SW ON light, located adjacent to the ac BUS TIE switch, illuminates when the ac BUS TIE switch is in

the ON position. The light receives 28-Vdc power from the essential dc bus through the GEN OUT LIGHT APU

circuit breaker on the copilot lower circuit breaker panel.

2.7.5 Secondary Ac System

The secondary ac power is comprised of two systems: the copilot ac instrument system, and the ac instrument and

engine fuel control system (see Figures 2-29 and 2-30).

2.7.5.1 Copilot Ac Instrument Power System

On aircraft prior to 165313, a single 250-volt-ampere inverter supplies 115-volt, 400-Hz, single-phase power. The

inverter draws dc power from the isolated bus; therefore, it can be operated from the battery during emergency

conditions of flight. Power can also be supplied from the essential ac bus to operate the pilot and copilot ac

instruments.

On aircraft 165313 and up, a 250-volt-ampere solid-state inverter supplies 115-volt, 400 Hz, single-phase ac power.

The inverters are operated by dc power from the isolated bus, and therefore, can be operated from the battery during

emergency conditions. During normal operations, power is supplied from the essential avionics ac bus to operate the

pilot and copilot instruments.

2.7.5.2 Ac Instrument and Engine Fuel Control System

On aircraft prior to 165313, the ac instruments and engine fuel control system are powered by a 115-volt, 400-Hz,

single-phase ac bus. One source of power is a 2,500-volt-ampere, single-phase inverter powered from the essential

dcbus through acurrent limiterlocated in theright-hand distribution box. Invertercontrol poweris provided through

the AC INST & ENG FUEL CONT INVERTER circuit breaker on the copilot lower circuit breaker panel from the

essential dc bus. Power can also be supplied from phase A of the essential ac bus through the AC INST & ENG FUEL

CONT PWR circuit breaker on the pilot side circuit breaker panel. The ac instruments and engine fuel control bus

circuits are protected from voltage spiking by a voltage surge suppressor (varistor) through the SURGE SUPPR

PROT DEV fuse on the pilot lower circuit breaker panel. Two instrument transformers are powered from the

115-volt, 400-Hz, single-phase bus, and provide 26-volt, single-phase, ac power for instrument use.

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On aircraft 165313 and up, the ac instruments and engine fuel control system are powered by a 115-volt, 400-Hz,

single-phase, ac instrument and engine fuel control bus. Power for the bus is supplied from one of two sources. One

source of power is a solid-state 1,000-volt-ampere inverter. Inverter control power is provided through the AC INST

& ENG FUEL CONT INVERTER circuit breaker on the copilot lower circuit breaker panel from the essential dc

bus. Power can also be supplied from Phase A of the essential avionics bus through the AC INST & ENG FUEL

CONT PWR circuit breaker on the pilot side circuit breaker panel. The ac instruments and engine fuel control bus

circuits are protected from voltage spiking by a voltage surge suppressor (varistor) through the SURGE SUPPR

PROT DEV fuse on the pilot lower circuit breaker panel. Two instrument transformers are powered from the

115-volt, 400-Hz, single-phase bus and provide 26-volt, single-phase, ac power for instrument use.

2.7.5.3 Secondary Ac System Controls

Controls for the secondary ac power system are located on the overhead electrical control panel (see Figure 2-28).

2.7.5.4 Copilot Ac Instruments Switch

The COPILOT AC INST switch is a three-position (STANDBY, OFF, NORMAL) rotary switch. In the STANDBY

position, poweris routed from theisolated dcbus to operate thecopilot instrument inverter forthe copilot instrument

powersupply system. IntheNORMALposition, theinverteristurned offand powerforthecopilot instrumentpower

system is taken from the essential ac bus. In the OFF position, no power is supplied to the system.

On aircraft 165313 and up, the COPILOTS AC INSTR switch is a three-position, rotary-type switch. In the STBY

(horizontal) position, dc power is routed from the isolated dc bus to operate the copilot instrument inverter for the

copilot instrument power supply system. In the ESS AV AC BUS position, the inverter is turned off and power for

the copilot instrument power supply system is supplied from the essential avionics ac bus. In the OFF position, no

power is supplied to the system.

2.7.5.5 Ac Instrument and Engine Switch

The AC INST & ENG switch is a three-position rotary switch. In the horizontal position, power is supplied to the

115-volt, 400-Hz, single-phase bus from phase A of the essential ac bus. In the vertical position, power is supplied

to theacinstruments and enginefuel control inverter, which will then powerthesystem. Iftheinverteroutputvoltage

is insufficient, the power supply is automatically switched from the essential dc bus to the essential ac bus. In the

OFF position, no power is supplied to the system.

On aircraft 165313 and up, the AC INSTR & ENG switch is a three-position rotary switch. In the NORM (ESS AV

AC BUS) position, power is supplied to the ac instrument and engine fuel control bus from phase A of the essential

avionics ac bus. In the STBY (horizontal) position, essential dc bus power energizes the ac instruments and engine

fuel inverter, which will then power the bus. If the inverter output voltage is insufficient, the power supply is

automatically switched from the inverter to the essential ac bus. In the OFF position, no power is supplied to the

system.

2.7.5.6 VOLTAGE & FREQUENCY SELECTOR Switch

The VOLTAGE & FREQUENCY SELECTOR switch has seven positions for measuring the output voltage and

frequency of the ac power supply sources. Placing the switch in the COPLT INV φA AC INST & ENG FUEL CONT

INV φC position, while simultaneously placing the PHASE SELECTOR switch in the PHASE A position, provides

an indication of the frequency and voltage of the copilot inverter on the frequency meter and the ac voltmeter,

respectively. Positioning the PHASE SELECTOR switch to PHASE C provides an indication of the output frequency

and voltage of the ac instrument and engine fuel control inverter. If the switch is at the COPLT INV φA AC INST

& ENG FUEL CONT INV φC position and essential ac power is being used in place of the inverter, the frequency

meter and the ac voltmeter will not indicate output frequency.

2.7.5.7 PHASE SELECTOR Switch

A three-position PHASE SELECTOR switch permits selection of the appropriate phase of electrical power when

measuring the output voltage and frequency of either of the inverters.

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2.7.6 Secondary Ac System Indicators

Indicators for the secondary ac power system are located on the overhead electrical control panel (see Figure 2-28).

2.7.6.1 Ac Voltmeter

An ac voltmeter permits measuring the output voltage of that phase of inverter power selected with the PHASE

SELECTOR switch. In order for the voltmeter to measureinverter output voltage, theVOLTAGE & FREQUENCY

SELECTOR switch must be in the COPLT INV φA AC INST & ENG FUEL CONT INV φC position. If the switch

is at either position and a bus source of power is being used in place of the inverter, however, the voltmeter will not

indicate output voltage.

2.7.6.2 Frequency Meter

A frequency meter permits measuring the frequency of the output power of that phase of inverter output selected with

the PHASE SELECTOR switch. In order for the frequency meter to measure the frequency of the inverter output

power, the VOLTAGE & FREQUENCY SELECTOR switch must be in the COPLT INV φA AC INST & ENG

FUEL CONT INV φC position. If the switch is at either position and a bus source of power is being used in place

of the inverter, however, the frequency meter will not indicate output frequency.

2.7.6.3 Selected Power OUT Lights

Two SEL PWR OUT press-to-test lights are located on the electrical control panel. If the COPILOT AC

INSTRUMENT or the AC INST & ENG fuel control SEL PWR OUT light illuminates, it indicates that no power

is available to therespective bus. When theAC INST & ENG switch is in theSTBY position and its SEL PWR OUT

light illuminates, the inverter has failed; however, the 115-Vac instrument and engine bus is then automatically

connected to the NORM power source. The light does not illuminate when the corresponding selector switch is OFF.

2.7.7 Dc Power System

Power from the essential ac bus and the main ac bus operates four transformer-rectifier units (two from each ac bus)

to provide dc power to the respective dc buses for the aircraft (see Figure 2-31). The four transformer-rectifier units,

mounted on theelectroniccontrol and supply rack, convert the powerfrom theac buses to 28-Vdc. Both theessential

ac bus and the main ac bus may be powered by any of the engine-driven generators (refer to Figure 2-32).

The essential ac bus is powered from the APU generator also, so it may be used as a source of dc power for ground

operation. The transformer-rectifier units feed current through reverse-current relays to the main dc bus and the

essential dc bus.

2.7.7.1 Dc System Buses

There are four buses in the DC POWER system: essential dc bus, battery bus, main dc bus, and isolated dc bus (see

Figure 2-33). The main and essential buses are connected through a reverse-current relay, which, in flight, allows

current to flow from the main bus to the essential bus but limits current flow in the opposite direction. When the

aircraft is on the ground, a touchdown switch is actuated to complete a circuit that overrides the reverse-current

limiting features of the reverse-current relay and permits current flow in either direction between the main and

essential buses. The essential and isolated buses are similarly connected through another reverse-current relay that

limits current flow from the isolated bus to the essential bus in flight. When the aircraft is on the ground, the

touchdown switch completes a circuit so that manual positioning of the dc BUS TIE switch overrides the

reverse-current limiting features of the reverse-current relay and permits current flow in either direction between the

isolated and essential buses.

Note

The dc BUS TIE switch is only effective if the touchdown switch is

actuated by the aircraft being on the ground.

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ENGINE DRIVEN GENERATORS

AC GENERATOR POWER SOURCE

AC BUS POWER SOURCE 2

ESSENTIAL

MAIN

AVIONICS

BUS

ESSENTIAL

MAIN

NO.

AC BUS

PRI

ALT

PRI

ALT

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

ESS

MAIN

APU Gen

ESS

Ext

Generator Out

Generator On

Note

The APU generator will power the main ac bus when the ac bus tie switch is placed in the on position. This function is

for ground checks only.

Aircraft 165313 and up.

If the ac bus tie switch is on, BSU no. 2 will select the main ac bus to power the main avionics bus.

Example:

No. 2 and No. 3 engine-driven generators out.

LH ac bus supplied by No. 1 generator.

Essential ac bus supplied by No. 1 generator.

Main ac bus supplied by No. 4 generator.

RH ac bus supplied by No. 4 generator.

If generator frequency drops below 368 Hz, the respective bus should transfer automatically.

Figure 2-32. Ac Bus Power Sources

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C-130T. FLIGHT MANUAL (2006) - page 2