C-141B COMMAND AIRCRAFT SYSTEMS TRAINING (CAST) …

C-141BCOMMAND AIRCRAFTSYSTEMS TRAINING

(CAST)

Program No. 5361-63

1 November 1994

Headquarters Air Mobility CommandMaintenance Management and Training

Scott AFB, IL

The C-141B Command Aircraft Systems Training (CAST) was produced by HQAMC/LGQRT at Scott AFB, IL. If you have questions or comments, please complete thecritique at the end of this training document or contact us through the following location.To request copies of this CAST, contact your training management section. If you should

have problems obtaining the information, contact:

HQ AMC/LGQRT402 Scott Drive

Unit 2A2Scott AFB, IL 62225-5308

DSN 576-4787FAX 576-1802

Direct Telecon is Authorized

C-141B CAST FOREWORDAir Mobility Command (AMC) demands and receivesthe very best from its maintenance community. Whetheraligned under operations, logistics or en route units, theaircraft maintenance mission is in the hands of reliableexperts who understand and proclaim “Global Reach,Global Power." As such, you are among those of anhonorable profession. You too have been chosen tocarry out our vital mission. For assuredly, withoutquality maintenance there can be no “Mobility.”

It is with this brief introduction that we welcome you tothis AMC Command Aircraft Systems Training (CAST)course. We believe that with the knowledge provided bythis weapon system training course, you will advancefurther ahead in your technical, supervisory, andmanagement development.

Although the Division for Logistics Studies cannot issuedegrees, we are deeply committed to weaponsystem-specific specialized training for advancing thetechnical expertise of all aircraft maintenance personnel.

As you go through this course, please remember that weare a significant part of the modernization programbeing implemented throughout the Department ofDefense (DOD). To maintain global superiority, wemust all expand the technological gap between ourselvesand our adversaries. That is your challenge. We wishyou the best in the complex and challenging world ofAMC aircraft maintenance.

iii

INTRODUCTIONThis AMC course is designed for 21XX aircraftmaintenance officers and 2AXXX career airmen. Itprovides a general working knowledge of the varioussystems, subsystems, and components installed on theC-141B aircraft. The course consists of three programs.You will be required to pass each end-of-program test tocomplete the course. The three programs are as follows:

PROGRAM TITLE

5361 General AircraftInformation

5362 Aircraft Systemsand Powerplant

5363 Avionics Systems

Finally, after satisfactorily completing the controlledtest for each of the three programs, you will be requestedto submit a critique through your local trainingmanagement center to us. We cannot overemphasize the

importance of your critique in improving this product.Every effort has and will be made to keep this coursecurrent and relevant to the overall goal of providingmaintenance people training to help perform theirassigned duties. Review conferences with subject matterexperts (SMEs) and a constant review of controlled testresults and critiques enable us to update and improve thiscourse. Comments, corrections, and/or questionsregarding course material or content are welcomed.These should be submitted to:

HQ AMC/LGQRT402 Scott Dirve

Unit 2A2Scott AFB, IL 62225-53084

Or call us direct at DSN 576-4787. We in the LogisticsBranch look forward to your successful completion ofthis systems training course and eagerly await sendingyou a diploma.

iv

GENERAL AIRCRAFT OVERVIEW

C-141B General InformationThe Lockheed C-141B Starlifter is a modern jet aircraft,designed primarily for transporting cargo. Powered byfour Pratt and Whitney TF33 turbofan engines, at 20,250pounds of thrust each, the aircraft can transportapproximately 70,000 pounds of cargo. The aircraft is168 feet 4 inches long with a 160-foot wing span. Thecargo compartment may be loaded with thirteen pallets,or configured to accommodate aft-facing seats,side-facing seats, or aeromedical evacuation of litterpatients.

Design features include a fully pressurized and airconditioned flight station and cargo compartment. Cargoloading is straight in from the rear, over an adjustableramp. Personnel loading is through troop doors on eachside of the fuselage aft of the center wing section or overthe cargo ramp. The single high wing is fullycantilevered and swept back at a 25-degree angle. Withair refueling capability, the aircraft range is limited onlyby mechanical and human limitations.

A high “T” tail provides improved operatingcharacteristics and simplified cargo loading. The flightstation contains provisions for a normal crew and reliefcrew. Facilities include a crew lavatory and galley.

An auxiliary power unit (APU), mounted in the left maingear pod, furnishes air for the aircraft pneumatic systemsand drives an AC generator to supply an alternate sourceof electrical power. The APU is operational only on theground and allows the aircraft to operate independent ofground support equipment when necessary.

Conventional, fully powered controls provide aircraftmaneuverability while airborne. Control about the rollaxis is provided by ailerons mounted on the outboardtrailing edge of each wing. Primary and backup for theailerons is supplied by aircraft hydraulic systems.Emergency operation is possible by mechanicallyoperated trim tabs, which are part of the ailerons. Controlabout the yaw axis is by a rudder connected to the trailingedge of the vertical fin, powered by aircraft hydraulicsystems. Control of the pitch axis is by an elevatormounted on the trailing edge of the horizontal stabilizer,which also is powered by the aircraft hydraulic systems.

Fowler-type wing flaps on the wing trailing edge, andspoilers mounted on upper and lower wing surface of

each wing serve to decrease aircraft speed and increasethe angle of descent. On the ground these units assist thewheel brakes and thrust reversers in minimizing groundroll.

All weather flying capability is ensured by wing andengine anti-ice, horizontal stabilizer de-ice, windshieldheat, and rain removal provisions. In addition to acomplex avionics system, the aircraft is equipped withan all-weather landing system (AWLS) that permitslanding with extremely limited visibility.

The C-141B Starlifter is a long-range, high-speed,high-altitude, swept wing monoplane, designed for useas a heavy logistic transport.

The designed ramp gross weight of the aircraft isnormally 325,000 pounds. The C-141’s value as astrategic cargo transport is its design for straight-in aftloading. Also, by design, the large unobstructed cargocompartment was built to be fully compatible with theAir Force 463L materials handling system.

Up to 13 standard 463L pallets may be loaded quickly.Alternate configurations will accommodate 166 troopsin aft-facing seats, 200 troops or 155 paratroops inside-facing seats, or 103 litter patients, and 14 additionalseats for attendants or ambulatory patients.

NOTE: Aircraft limited to 200 troops due to presentoxygen system.

Aircraft SystemsThe following are a brief discription of the systems onthe C-141B aircraft. Each will be explained later in theappropriate CAST volume.

Power PlantThe TF33-P-7A power plant has a sixteen-stageaxial-flow split compressor; an eight-can, can-annularcombustion chamber; and a four-stage axial-flow typeturbine. There are two accessory cases on the engine.

The engine is divided into five operating sections:compressor (including the fan section), diffuser section,combustion section, turbine exhaust, and accessorydrive sections.

v

Fuel SystemThis is a ten-tank, wet-wing, integral manifold fuelsystem. The four main tanks, four auxiliary tanks, andtwo extended range tank hold 153,352 pounds of usablefuel. The fuel system is capable of supplying any enginefrom any tank, transferring fuel from any tank to anyother tank in flight or on the ground, air refueling, singlepoint refueling, and jettisoning.

Electrical SystemThe C-141B has a parallel AC and DC power system.The system is designed so that:

(1) No single failure or probable combination of failuresshall cause complete loss of electrical power.

(2) System operation and protection shall be asautomatic as practical.

(3) External power shall not be required for normalengine start.

The C-141B uses a 200/115 volt, 400 hertz, three-phaseAC power system as the primary source of electricalpower.

In flight, this power is supplied by four 50-kilovoltampere (KVA) engine-driven generators. Duringground operation, power is supplied by one of threesources: an auxiliary power unit (APU) drivengenerator, an external power source, or theengine-driven generators.

During inflight emergencies, power is supplied by a twoKVA hydraulically driven emergency generator. Thefour engine-driven generators, the APU generator, andexternal power are controlled from the flight engineer’spanel. Emergency generator operation is automatic, butcan be controlled manually from the pilot’s instrumentpanel.

Hydraulic SystemsThe C-141B hydraulic system consists of separate andfunctionally independent systems designated as systemsNo. 1, 2, 3, and an emergency nose landing gearextension system. Each system is divided into ahydraulic power system and system components towhich pressure is delivered. MIL-H-83282 type fluid isused. Each of the hydraulic systems has a service centerwhere the hydraulic reservoirs are located. The hydraulicsystems controls and indicators are located on the lowerleft corner of the flight engineer’s upper panel and pilotsoverhead control panel.

Landing GearThe fully retractable, tricycle landing gear consists ofdual nose wheel assembly mounted under the forwardfuselage and two dual-tandem main gear assembliesmounted in pods attached to each side of the fuselage.Deceleration on the ground is accomplished by eightmultiple disc-type wheel brakes with full anti-skidprotection and reverse thrust provisions on each of thefour engines.

Air Conditioning SystemsThe air used for air conditioning is supplied by the bleedair manifold, regulated to a maximum pressure of 70 psiby the system air pressure regulating and shutoff valves.Airflow from the left air conditioning pack normallysupplies one third of its flow to the flight station and theremainder to the cargo compartment. Airflow from theright pack is normally routed to the cargo compartment.

Air conditioning is accomplished by cooling bleed

manifold air to 230o

C through a primary heat

exchanger, then to 65oC through a secondary heat

exchanger, and routing a portion of air through a

turbine refrigeration unit which super cools the air. To

prevent ice formation, a certain amount of 65oC air is

remixed with super cooled air to maintain a

temperature of 20o

C downstream of the water

separator. The 20oC air can then be mixed with an

appropriate amount of 230oC air to obtain the desired

cabin temperature.

The two identical air conditioning packs are located inthe center wing section and are made up of the followingsubunits: Primary heat exchanger, secondary heatexchanger, refrigeration unit, distribution ducting,temperature control system, flow control and shutoffvalve.

Flight Controls SystemsThe aircraft is controlled by hydraulically poweredaileron, rudder, and elevator systems. Aerodynamiclateral control is available through a cable-controlledaileron tab, if normal hydraulic pressure is lost.Electrically operated trim systems are provided fortrimming the aircraft about the roll, yaw, and pitch axis.A manually operated, hydraulically powered trimsystem and an electrically operated, hydraulicallypowered trim system are also provided for the pitch axis.

The primary flight control power control assembly hasanticavitation valves to prevent cavitation of theactuators when they are operating with hydraulic poweroff, while the shutoff and bypass valve is in the normalposition (OPEN).

vi

Oxygen SystemThe aircraft is equipped with two independent liquidoxygen systems. One for the crew which suppliesoxygen to crew members for normal and emergencyoperations. The second one is for personnel in the cargocompartment, both crew and passengers .

Auxiliary Power Unit (APU)The APU, located in the left wheel well, supplies air forengine starting, air for the environmental systems, andmechanically drives an AC generator. The APUgenerator is identical to the main generators, and is forground operation only. The APU controls receive powerfrom the isolated DC bus through circuit breakers on theflight engineer’s No. 3 circuit breaker panel.

AvionicsThe avionics systems installed on the C-141B aircraftare grouped under several major categories. Thesedifferent categories are based on the primary function agiven avionics system performs. For our purpose, wehave grouped together all of the “similar” systems ineach section of program 5363. Some of the systems,however, stand alone even though they either impactgreatly the performance of, or depend on other systeminputs/outputs to perform their primary functions.

CommunicationCommunications comprise the systems and componentswhich provide a means of communication from onestation to another within the aircraft, and between theaircraft and other aircraft or ground stations. Thesesystems are radio, interphone/intercom, satellitecommunication (SATCOM), passenger address, voicerecorder, and precipitation static discharging. Therespective systems may operate independently, orinterface with others to satisfy operational requirements.Communication components are located in the flightcompartment, passenger compartment, avionicscompartment, inside the T-tail section, and at thosemaintenance points where ready communication isrequired.

Identification Friend or Foe (IFF)The IFF system is the military version of the air trafficcontrol radar beacon system (ATCRBS, also calledATC). The air traffic control (ATC) transponder systemprovides a ground based control center with aircraftlocation and identification. The system provides theground control with a series of pulses (response) from

the aircraft to the ground station when triggered by aseries of pulses (interrogation) from the ground station.The IFF transponder responds to interrogating pulses.

CompassThe compass systems supplies the pilot and copilot withdirectional data used in navigating the aircraft. Themagnetic information supplied by the compass system isdisplayed on the applicable horizontal situationindicator (HSI). Output signals from the compasssystem are provided to other systems in the airplanerequiring a heading reference.

InstrumentsThe instrument system is used to give readings of manydifferent system operations. Each system has its owninstruments designed to indicate that system’soperation. These instruments will be explained in thesection covering that system.

Inertial Navigation Systems (INS)The INS is an all weather, worldwide inertial navigationsystem with automatic radio position update capability.The system may be operated as an area navigationsystem or as an inertial navigation system. Whenoperated as an area navigation system, using range andbearing data from selected TACAN stations, the systemprovides accurate horizontal navigation independent oftime and inertial navigation system drift errors. Thesystem also provide an orbit and rendezvous mode tosteer the aircraft to a predetermined orbit pattern. Thereare three inertial navigation systems installed in theaircraft.

Auto FlightThe auto flight system consist of units and componentsthat furnish a means of automatically controlling theflight path, safe reference speed, and trim condition ofthe aircraft. Auto flight consist of four subsystems:autopilot, flight director, automatic pitch trim (APT),and autothrottle/speed control (AT/SC).

Fire ProtectionFire protection consist of a fire detection system, a

smoke detection system, and a fire extinguishing

system. The fire detection system provides the means to

detect a fire or overheat condition and to alert the crew

by aural and visual indications. The smoke detection

system monitors the cabin and cargo area and causes

a caution indicator to come on in the fl ight

vii

compartment. The extinguishing system provides the

means to extinguish the fire.

VI

GLOSSARY OF ABBREVIATIONS/ACRONYMSThe following list of terms was taken from all the CAST books presently in use. Not all the terms in this list will pertainto your particular weapon system.

AA/A Air-to-Air

ABS Asymmetry Brake System(Flaps And Slats)

AC; ac Alternating Current

ACC Avionics Computer Control

ACFT Aircraft

ACM Air Combat Mode

ACP Automatic CommunicationProcessor

AC&PS Air Conditioning &Pressurization System

ACU Avionics Control Unit

ADF Automatic Direction Finder

ADG Accessory Drive Gearbox

ADI Attitude Director Indicator

ADS Air Data Sensor; Air DeliverySystem

ADSP Analog Data Signal Processor

AFCS Automatic Flight ControlSystem

AFGS Automatic Flight GuidanceSystem

AFSATCOM Air Force SatelliteCommunications

AGE Aerospace Ground Equipment

AGM Air-to-Ground Missile

AHRS Attitude Heading ReferenceSystem

AIL Aileron

AIL STA Aileron Station

AILA Airborne Instrument LandingApproach

AIMS Air Traffic Control RadarSystem

ALT Altitude; Altimeter; Alternate;Alternator

ALDCS Air Lift Distrubution ControlSystem

AM Amplitude Modulation

AMI Airspeed Mach Indicator

AMPL Amplifier

AMUX AvionicsMultiplex

ANT Antenna

AOA Angle-of-Attack

AOAT Angle-of-Attack Transmitter

AP Auto Pilot; Array Processor

APT Automatic Pitch Trim

APU Auxiliary Power Unit

ATM Air Turbine Motor

AR Air Refuel

ARB Aerial Refuel Boom

ARO Aerial Refueling Operator

A-1

ASSY Assembly

ATCS Automatic Thrust ControlSystem

ATF Automatic Terrain Following

ATS Air Turbine Starter; AutoThrottle System

ATT Attitude

AU Accelerometer Unit

Auto Automatic

AUX Auxiliary

AVVI Altitude Vertical VelocityIndicator

BBAL Balance

BALU Bus Allocation Logic Unit

BARO Barometric

BCN Beacon

BCU Bus Control Unit

BIT Built-in Test

BITE Built-in Test Equipment

BL Buttock Line

BPP Bus Protection Panel

BRG Bearing

BRT Bright; Brightness

BSC Beam Steering Controller

BTB Bus Tie Breaker

BTC Bus Tie Contacter

BU Battery unit

BW Bandwidth

BPO Basic Post Flight

CC Celsius; Centigrade

CADC Central Air Data Computer

CAL Calibration

CAM Content-Addressable Memory

CAMS Core Automated MaintenanceSystem

CAP Capture

CARA Combined Altitude RadarAltimeter

CAS Computed Airspeed

CASS Centralized Aircraft ServicingSystem

CAST Command Aircraft SystemsTraining

CB Circuit Breaker;Bromocloromethane

CC Cubic Centimeter

CCP Computer Control Panel

CCU Central Control Unit;Compass Compensation Unit

CCW Counter-Clockwise

CDI Course Deviation Indicator

CDIU Controls and DisplaysInterface Unit

CDPIR Crash Data Position IndicatorRecorder

CDU Controls and Displays Unit

CG Center of Gravity

CIT Compressor InletTemperature

CITS Central Intergrated TestSystem

A-2

CLR Clear

CMD Command

COMBS Contractor Operated andMaintained Base Supply

COMP Compass

CONT Control; Continuious

CO2 Carbon Dioxide

CP Command Post

CPLR Coupler

CPS Cycles Per Second (Hertz)

CPU Central Processing Unit

C/RAD Command Radio

CRS Course

CRT Cathode Ray Tube

CSD Constant Speed Drive

CTC Communication and TrafficControl

CU Cubic

CW Clockwise; Continuious Wave

CWS Control Wheel Steering

DDB Decibel(s)

DC; dc Direct Current

DCU Display Control Unit

DEST Destination

DF Direction Finder

DFDR Digital Flight Data Recorder

DIA Diameter

DICU Display Interface Control Unit

DIFF Differential

DME Distance MeasuringEquipment

DP Differential Protection

EEBADS Engine Bleed Air Distribution

System

ECM Electronic Countermeasures

ECS Environmental ControlSystem

EGT Exhaust Gas Temperature

EGCU Emergency Generator ControlUnit

ELEV Elevator

ELEV STA Elevator Station

EMER Emergency

EPD Electrical Power Distribrution

EPR Engine Pressure Ratio

ET Elapsed Time

ETA Estimated Time of Arrival

EXT External

FF Fahrenheit

F&LC Frequency and LoadController

FCI Flight Command Indicator

FCP Flight Computer Program

FD Flight Director

FDC Flight Director Computer

FDS Flight Director System

A-3

FL Flight Level

FM Frequency Modulation

FMC Fully Mission Capeable

FMP Fuel Management Panel

FOL Foward Operating Location

FREQ Frequency

FR BM STA Front Beam Station

FRZE Freeze

FS Fuselage Station

FSAS Fuel Savings AdvisorySystem

FSS Fire Suppresion System

FT; ft Foot; Feet

FWD Forward

GG Unit of Gravity; Gram(s)

GA General Aircraft

GAAS Go Around Attitude System

GCU Generator Control Unit

GHz Gigahertz

GMB Ground Marker Beacon

GMT Greenwich Mean Time

GOX Gaseous Oxygen

GPM Gallons Per Minute

GPWS Ground Proximity WarningSystem

GS Glideslope; Groundspeed

GSE Ground Support Equipment

GSU Ground Servicing Unit (LN2Servicing Truck)

GYRO Gyroscope

HHDG Heading

HF High Frequency

HG Inches of Mercury

HOR Horizontal

HPO Hourly Post Flight

HSI Horizontal Situation Indicator

HSC Home Station Check

HZ Hertz

IIAS Indicated Airspeed

ICS Intercom System

IDENT Identification of Position(IFF)

IF Intermediate Frequency

IFF Identification Friend or Foe

IFR Inflight Refuel

IGV Inlet Guide Vane

ILS Instrument Landing System

IN Inches

INBRD Inboard

IND Indicator

INS Inertial Navigation Syatem

INST Instrument

INT Internal; Intensity

INTER Internal

INU Inertial Navigation Unit

A-4

ISOL Isolation

JJG Job Guide

JOAP Joint Oil Analysis Program

JTSN Jettison

KKHz Kilohertz

Kv Kilovolts

KVA Kilovolt/Ampere

KW Kilowatt

LLAT Latitude

LB Pounds

LBL Left Buttock Line

LH Left Hand

LIM Limit

LIN Linear

LN2 Liquid Nitrogen

LOC Localizer

LOG Logarithmic

LOX Liquid Oxygen

LRU Line Replaceable Unit

LSB Lower Sideband

LVDT Linear Variable DifferentialTransformer

MMAC Mean Aerodynamic Chord

MADAR Malfunction DetectionAnalysis and Recording

MAG Magnetic

MALF Malfunction

MAN Manual

MAX Maximum

MHz Megahertz

MIC Microphone

MILSPEC Military Specification

MKR Marker; Marker Beacon

MLG Main Landing Gear

MM Millimeter

MMR Multi-Mode Radar

MOB Maintenance Operating Base

MOI Maintenance OperatingInstruction

MON Monitor

MPH Miles Per Hour

MS Maintenance Support

MSG Message

MUX Multiplexer

MWCS Master Warning and Caution

NNAV Navigator; Navigation

NC Core Engine Speed

NF Fan Engine Speed

NICAD Nickel-Cadmium

NLG Nose Landing Gear

NM Nautical Miles

A-5

NO Number

NORM Normal

NWS Nose Wheel Steering

NSN National Stock Number

OOPER Operate

OVRD Override

OUTBRD Outboard

PPA Public Address

PACS Pilot Assist Cable System

PCA Power Control Assembly

P&D Pressurization and Dump

PGM Program

POS Position

PPH Pounds Per Hour

PP STA Power Plant Station

PP WL Power Plant Water Line

PS Power Supply

PSF Pounds Per Square Foot

PSI Pounds Per Square Inch

PSIA Pounds Per Square Inch(Absolute)

PSID Pounds Per Square Inch(Differental)

PSIG Pounds Per Square Inch(Guage)

PTO Power Take off (ADG)

PTT Press-To-Test

PTU Power Transfer Unit

PWR Power

QQAD Quick Attach/Detach

QD Quick Disconnect

QEC Quick Engine Change

QTY Quanity

QUAL Quality

RRA Radar Altimeter

RADAR Radio Detection And Ranging

RAI Radar Altimeter Indicator

RCVR Receiver

RDR Radar

REC Receive

REL Relative

REV Reverse

RF Radio Frequency

RG Range

RH Right Hand

RPM Revolutions Per Minute

RT Receiver Transmitter

RTN Return

RTV Rubber Temperature

RUD Rudder

SSADI Standby Attitude Director

Indicator

A-6

SAR Signal Acquistion RemoteUnit

SASS Stability AugmentationSub-System

SATCOM Sattellite Communication

SCAS Stability & ControlAugmentation System

SCM Space Cargo Modification

SE Support Equipment

SLV Slave

SMK Special Mission Kit

SOAP Spectrometric Oil AnalysisProgram

SPOT Spotlight

SPR Single Point Refuel(Receptacle)

SQ; SQL Squelch

SRU Shop Replaceable Unit

STAB Stabilizer; Stabalized

STBY Standby

SYNC Synchronization

TTACAN Tactical Air Navigation

TAS True Air Speed

TAT Total Air Temperature

TCTO Time Compliance TechnicalOrder

TD Time Delay

TEMP Temperature

TF Terrain Following

TFR Terrain Following Radar

TH True Heading

T.O. Technical Order

T/R Transmitt/Receive

TR Transformer Rectifier; ThrustReverser

TRT Takeoff Thrust

UUHF Ultra-High Frequency

USB Upper Side Band

VVAC; vac Volts Alternating Current

VDC; vdc Volts Direct Current

VHF Very High Frequency

VOL Volume

VOR VHF Omnirange; VisualOmnirange; VaribleOmnirange

VSFI Vertical Scale Flight Indicator

VSI Vertical Scale Indicator

VSV Variable Stator Vane

WWD Weather Detection

WOW Weight On Wheels

WS Wing Sweep

WL Water Line

XXMIT Transmit

XMTR Transmitter

A-7

C-141BPROGRAM 5361

GENERAL AIRCRAFT


Scott AFB, IL

PROGRAM 5361GENERAL AIRCRAFT

TABLE OF CONTENTS

SECTION ISTRUCTURAL DOORS

PARA TITLE PAGE

1-1 STRUCTURAL DOORS . . . . . . . . 1-3

1-2 CREW AND TROOP DOORS. . . . . . 1-3

1-3 EMERGENCY EXITS SYSTEM. . . . 1-4

1-4 CARGO DOORS SYSTEMS. . . . . . 1-10

Pressure Door. . . . . . . . . . . . . 1-10

Cargo Ramp. . . . . . . . . . . . . .1-10

Petal Doors . . . . . . . . . . . . . .1-10

1-5 INTERIOR DOORS SYSTEM. . . . . 1-14

1-6 DOOR WARNING SYSTEM. . . . . . 1-14

SECTION IIGROUND HANDLING, INSPECTION,

AND SERVICING

PARA TITLE PAGE

2-1 GROUND HANDLING, INSPECTION,AND SERVICING . . . . . . . . . . . . 2-3

2-2 MAINTENANCE GENERALINFORMATION . . . . . . . . . . . . . 2-3

General Maintenance Practices. . . . . 2-3

2-3 GROUND HANDLING . . . . . . . . . 2-3

Towing . . . . . . . . . . . . . . . . . 2-3

Parking . . . . . . . . . . . . . . . . . 2-4

Jacking . . . . . . . . . . . . . . . . . 2-4

Fuel System. . . . . . . . . . . . . . . 2-4

Hydraulic System . . . . . . . . . . . 2-4

Oxygen System. . . . . . . . . . . . . 2-4

2-4 INSPECTIONS. . . . . . . . . . . . . . 2-4

Aircraft Inspections Information. . . . 2-4

2-5 SERVICING . . . . . . . . . . . . . . . 2-5

Aircraft Servicing Information. . . . . 2-5

Fuel . . . . . . . . . . . . . . . . . . . 2-5

LOX Servicing . . . . . . . . . . . . . 2-5

Hydraulic . . . . . . . . . . . . . . . . 2-5

Nitrogen Servicing. . . . . . . . . . . 2-5

Potable Water. . . . . . . . . . . . . . 2-5

Engine Oil Service. . . . . . . . . . . 2-5

Lubrication . . . . . . . . . . . . . . . 2-5

SECTION IIIEQUPMENT AND FURNISHINGS

PARA TITLE PAGE

3-1 EQUIPMENT AND FURNISHINGS . . 3-3

3-2 FLIGHT STATION EQUIPMENT. . . . 3-3

Flight Station Seats. . . . . . . . . . . 3-3

Relief Crew Bunks. . . . . . . . . . . 3-3

iii

Side Consoles. . . . . . . . . . . . . . 3-3

Stowage Compartments. . . . . . . . . 3-4

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

Electrical Spares Box. . . . . . . . . . 3-4

Technical Order (T.O.) FileBookshelf. . . . . . . . . . . . . . . . 3-4

Miscellaneous Flight StationEquipment . . . . . . . . . . . . . . . 3-4

Flight Station Entrance Ladder. . . . . 3-4

Hand Rails and Assist Handles. . . . . 3-4

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

Hooks and Hangers. . . . . . . . . . . 3-4

Checklist and Letdown ChartHolders . . . . . . . . . . . . . . . . . 3-4

Glare Shield. . . . . . . . . . . . . . . 3-4

Sun Visors . . . . . . . . . . . . . . . 3-4

Cupholders. . . . . . . . . . . . . . . 3-9

Fuel Tank Dipstick. . . . . . . . . . . 3-9

3-3 GALLEY . . . . . . . . . . . . . . . . . 3-9

General Galley Provisions. . . . . . . 3-9

Fresh Water Tanks. . . . . . . . . . . 3-9

Refrigerator . . . . . . . . . . . . . .3-10

Oven. . . . . . . . . . . . . . . . . .3-10

Hot Cup . . . . . . . . . . . . . . . .3-10

Hot Beverage Unit. . . . . . . . . . . 3-10

Galley Control Panel. . . . . . . . . 3-10

Waste Container. . . . . . . . . . . . 3-10

3-4 LAVATORY . . . . . . . . . . . . . .3-10

Toilet. . . . . . . . . . . . . . . . . .3-10

Wash Water. . . . . . . . . . . . . .3-10

Waste Water Tank. . . . . . . . . . . 3-10

Waste Paper Container. . . . . . . . 3-10

Servicing Provisions. . . . . . . . . . 3-10

Miscellaneous Equipment. . . . . . . 3-10

3-5 CARGO COMPARTMENTEQUIPMENT . . . . . . . . . . . . . .3-11

3-6 EMERGENCY EQUIPMENT . . . . . 3-12

Escape Ladders. . . . . . . . . . . . 3-12

Escape Ropes. . . . . . . . . . . . . 3-12

Life Vests . . . . . . . . . . . . . . .3-12

First-Aid Kits . . . . . . . . . . . . . 3-12

Fire Extinguishers. . . . . . . . . . . 3-12

Crash Ax . . . . . . . . . . . . . . .3-12

Cargo Compartment Warning Horn . . 3-12

3-7 AERIAL DELIVERY SYSTEM(ADS) . . . . . . . . . . . . . . . . . .3-12

3-8 SPECIAL MISSIONS EQUIPMENT . . 3-14

Oxygen Distribution and Supply Kit . . 3-14

Life Raft Liner Kit . . . . . . . . . . . . 3-14

Stanchion Kit . . . . . . . . . . . . . .3-14

Comfort Pallet. . . . . . . . . . . . . .3-14

Litter Configuration. . . . . . . . . . . 3-14

Troop Seats. . . . . . . . . . . . . . .3-14

Paratroop Seat Configuration. . . . . . 3-14

Paratroop Spoiler Doors. . . . . . . . . 3-14

Anchor Cables. . . . . . . . . . . . . .3-14

Portable Urinal (Honey Bucket). . . . . 3-14

iv

SECTION IVLIGHTING SYSTEM

PARA TITLE PAGE

4-1 LIGHTING SYSTEM . . . . . . . . . . 4-3

4-2 FLIGHT STATION LIGHTINGSYSTEM . . . . . . . . . . . . . . . . . 4-3

4-3 CARGO AND SERVICE COMPART-MENT LIGHTING SYSTEM . . . . . . 4-3

Cargo Compartment Dome Lights. . . 4-4

Cargo Compartment Overhead Lights . 4-4

Crew Rest Platform. . . . . . . . . . . 4-5

Lavatory Light . . . . . . . . . . . . . 4-5

Flight Station Access Light . . . . . . 4-6

Cargo Ramp Loading Lights. . . . . . 4-6

Nose and Main Wheel Well Lights. . . 4-7

Under Flight Deck Rack Lights. . . . . 4-7

Vertical Stabilizer Lights. . . . . . . . 4-7

Platform Lights. . . . . . . . . . . . . 4-7

4-4 EXTERIOR LIGHTING SYSTEM. . . . 4-7

Landing Lights . . . . . . . . . . . . . 4-7

Formation Lights. . . . . . . . . . . . 4-7

Navigation Lights . . . . . . . . . . . 4-7

Anti-Collision Lights . . . . . . . . . . 4-7

Taxi Lights . . . . . . . . . . . . . . . 4-7

Wing Leading Edge Lights. . . . . . . 4-7

Aerial Refueling Fairing Lights. . . . . 4-7

Aerial Refueling Slipway Lights. . . . 4-7

4-5 EMERGENCY LIGHTING SYSTEM . 4-10

SECTION VFIRE PROTECTION SYSTEM

PARA TITLE PAGE

5-1 FIRE PROTECTION SYSTEM. . . . . 5-3

5-2 DETECTION AND WARNINGSYSTEMS . . . . . . . . . . . . . . . . 5-5

5-3 FIRE EXTINGUISHING SYSTEMS . . 5-7

SECTION VILANDING GEAR SYSTEM

PARA TITLE PAGE

6-1 LANDING GEAR SYSTEM. . . . . . . 6-3

6-2 MAIN LANDING GEAR AND DOORSSYSTEM . . . . . . . . . . . . . . . . . 6-3

Shock Strut. . . . . . . . . . . . . . . 6-3

Axles . . . . . . . . . . . . . . . . . . 6-3

Wheels, Tires, and Brakes. . . . . . . 6-3

Doors . . . . . . . . . . . . . . . . . . 6-3

6-3 NOSE LANDING GEAR AND DOORSSYSTEM . . . . . . . . . . . . . . . . . 6-5

Doors . . . . . . . . . . . . . . . . . . 6-5

6-4 EXTENSION AND RETRACTIONSYSTEM . . . . . . . . . . . . . . . . . 6-8

Normal Extension and Retraction. . . 6-8

Emergency Extension. . . . . . . . . . 6-8

6-5 MAIN LANDING GEAR BRAKES ANDANTI-SKID SYSTEM . . . . . . . . . . 6-8

Brakes. . . . . . . . . . . . . . . . . . 6-8

Anti-Skid . . . . . . . . . . . . . . . . 6-8

6-6 NOSE LANDING GEAR STEERINGSYSTEM . . . . . . . . . . . . . . . .6-13

v

6-7 LANDING GEAR POSITION ANDWARNING SYSTEM. . . . . . . . . . 6-13

Nose Gear. . . . . . . . . . . . . . .6-13

Indication . . . . . . . . . . . . . . .6-13

Visual/Audio Warning. . . . . . . . . 6-14

SECTION VIIFLIGHT CONTROL SYSTEM

PARA TITLE PAGE

7-1 FLIGHT CONTROL SYSTEM . . . . . 7-3

7-2 PRIMARY FLIGHT CONTROLSYSTEM . . . . . . . . . . . . . . . . . 7-3

Control Cables. . . . . . . . . . . . . 7-4

Tension Regulators. . . . . . . . . . . 7-4

Artificial Feel . . . . . . . . . . . . . . 7-4

7-3 AILERON SYSTEM. . . . . . . . . . . 7-4

Servotabs. . . . . . . . . . . . . . . . 7-4

Aileron Trim System. . . . . . . . . . 7-4

7-4 RUDDER SYSTEM . . . . . . . . . . . 7-8

Power Control Unit HydraulicPressure. . . . . . . . . . . . . . . . . 7-8

Rudder Trim System. . . . . . . . . . 7-8

7-5 ELEVATOR SYSTEM. . . . . . . . . . 7-8

Artificial Feel System. . . . . . . . . 7-11

Loss of Normal Hydraulic Pressure . . 7-11

Stall Prevention System. . . . . . . . 7-11

7-6 SECONDARY FLIGHT CONTROLS . 7-12

Control Cables. . . . . . . . . . . . . 7-12

Tension Regulator. . . . . . . . . . . 7-12

7-7 PITCH TRIM SYSTEM. . . . . . . . . 7-13

Manual Hydraulic and Electro-Hydraulic Systems. . . . . . . . . . . 7-13

Electric System. . . . . . . . . . . . 7-13

Travel and Position. . . . . . . . . . 7-13

7-8 FLAP SYSTEM. . . . . . . . . . . . . 7-13

Flap Asymmetry System. . . . . . . 7-13

7-9 SPOILER SYSTEM. . . . . . . . . . . 7-14

Spoiler Asymmetry System. . . . . . . 7-16

Safety Features. . . . . . . . . . . . . 7-16

7-10 FLIGHT CONTROL HYDRAULICSYSTEMS. . . . . . . . . . . . . . . .7-21

No.1 Hydraulic System. . . . . . . . . 7-21



7-11 DEPLETED URANIUMCOUNTERWEIGHTS . . . . . . . . . 7-21

vi

SECTION I STRUCTURAL DOORS

This section consists of information extracted from the T.O. 1C-141B-2-52GS-00-1.

After reading this section, you should be able to recall the following:

1. The purpose of the structural doors

2. The function of

• crew and troop doors

• emergency exits

• cargo door

• interior doors

• door warning

3. The location of components for

• crew and troop doors

• emergency exits

• cargo door

• interior doors

• door warning

Structural Doors Systems

1-1

Command Aircraft Systems Training

1-2

1-1 STRUCTURAL DOORS

The structural doors system of the C-141B aircraftincludes the crew and troop doors, emergency exits,cargo doors, interior doors, and door warning. Theoverall functions and operation of the structural doors

system are described in this section. A more detaileddescription of each subsystem is presented in subsequentsections of this program.

1-2 CREW AND TROOP DOORS SYSTEM

The crew and troop doors provide entry to and exit fromthe aircraft . The crew entrance door, which is located onthe left side of the cargo compartment just aft of theforward bulkhead, is normally used for entrance and exitfrom the aircraft (Fig 1-1). The troop doors, located oneach side of the cargo compartment just forward of theramp, are normally used for exit of paratroop personnel,although they may be used during maintenance (Fig1-2). Each door has a combination of control rods andbellcranks used to position latches. These latches extend

from the doors to lock and retract into the door to unlock.A combined interior handle and exterior handle on eachdoor is used to control the position of the latches. Thecrew door uses one counterbalance assembly while eachtroop door uses two counterbalance assemblies. Thesecounterbalances are used to carry the weight of the doorto aid personnel when opening or closing the door. Acable attached between the counterbalance and the doorprovides the connection between these two components.


1-3

Fig 1-1, Crew entrance door

1-3 EMERGENCY EXITS SYSTEM

The emergency escape system provides escape from orentry into the aircraft during emergencies. There are twotypes of emergency exits: emergency escape doors andemergency escape hatches. Four emergency escapedoors are installed in the fuselage walls: two in theforward cargo compartment and two in the aft cargocompartment (Fig 1-3). The doors can be completelyremoved from the fuselage when the latches are released.The latches are controlled by a combination handle,control rod, and torque tube. There are four emergencyescape hatches: the flight station emergency escapehatch, the forward cargo compartment emergencyescape hatch, the left aft cargo compartment emergencyescape hatch, and the right aft cargo compartmentemergency escape hatch. The No. 1 hatch is locatedoverhead in the flight station and can be released fromthe inside or outside (Fig 1-4). The hatch, when released,opens down and into the flight station. The hatch hasquick-release hinges so it can be removed and set out ofthe way. The No. 2 hatch is located overhead just aft ofthe crew door and can be released from the inside or

outside (Fig 1-5). The hatch, when released, opens upand forward to the outside of the aircraft. The No. 3 hatchis located overhead just aft of the aft beam of the centerwing (Fig 1-6). This hatch can be released from theinside or outside. The hatch must be supported when thelatch is released since it is not connected or hinged to theaircraft structure. The No. 4 hatch is located above theleft troop door and can be opened from the inside oroutside (Fig 1-7). If opened from the inside, the hatchopens downward and into the cargo compartment. Ifopened from the outside, the hatch is totally releasedfrom the aircraft and falls into the cargo compartment.The three cargo compartment emergency escape hatcheshave ladders mounted nearby so personnel can reach thehatches. The three cargo compartment hatches and thefour escape doors have ropes mounted nearby to providea means of getting off the aircraft after exiting the hatchor door. Although not part of the emergency escapesystem, a stabilizer access door is installed at the top ofthe T- tail. This access door can be reached from the topof the forward ladder and is used to gain access to the topof the horizontal stabilizer.


1-4

Fig 1-2, Troop door


1-5

Fig 1-3, Emergency exit doors


1-6

Fig 1-4, No. 1 emergency escape hatch


1-7



1-8



1-9


1-4 CARGO DOORS SYSTEM

The clam shell-type petal doors, pressure door, and rampare referred to as the cargo doors. Hydraulic pressure foroperation of the doors is provided by hydraulic systemNo. 3. Four separate control panels may be used tooperate the cargo doors. Pilot’s and copilot’s aerialdelivery system (ADS) control panels permit operationof the cargo doors in flight. These panels also containindicator lights which visually display the position of thedoors. The ramp control panel, located near the ramp, isused to control the cargo doors on the ground. A switchon the crew door interphone and public address (PA)control panel is normally used to open or close thepressure door only in flight. When static line A-frameactuators are installed, only the PRESSURE DOORswitch on the ramp control panel will close the pressuredoor. If the static line A-frame actuators are not installed,dummy plugs (using jumpers in place of the actuator) areinstalled. When the dummy plugs are installed,operation of the pressure door is possible at any of thefour panels. The electrically-controlled cargo doorsmust open in a set sequence to prevent damage.

Pressure DoorThe pressure door seals off the cargo compartment fromthe unpressurized aft fuselage when closed (Fig 1-10). Itlocks to the ramp through a series of hooks, which rotatelocked or unlocked by a pressure door lock actuator ateach lower corner of the pressure door (Fig 1-10). Asingle push-pull type hydraulic actuator, mounted on theaft side of the pressure door, opens or closes the pressuredoor (Fig 1-11). The pressure door has a continuouspiano-type hinge at the top and opens up and aft (Fig1-9). An uplock in the aft upper deck mechanically locksthe pressure door open and swings a mechanicalindicator down for visual confirmation. This uplock isoperated by a hydraulic pressure door uplock actuator.During a normal opening sequence, the pressure doormoves forward and unlocks simultaneously. After thepressure door is against the ramp, it moves open into theuplock. During a normal close sequence, the pressuredoor moves up (open) to allow the uplock to release, thenlowers until it contacts the ramp. At this point, the hooksrotate locked and the pressure door moves aft until thehooks engage the ramp latch fittings. The system willremain stationary with the pressure door lock and opencircuits energized (called the holding circuit). This aidsin the installation of cam jacks to mechanically duplicatethe pressure door holding circuit (Fig 1-9). After theholding circuit is de-energized, seven auxiliary latch

assemblies can be installed as a backup to the primarylock system.

Cargo RampThe ramp provides truck-bed or ground-level loadingaccess to the cargo compartment. The ramp along withthe pressure door provides the aft closure for the cargocompartment . The ramp is locked to the aircraftstructure by a dead-bolt type locking system. A series ofconnecting rods and locking pins secure fittings on theouter edge of the ramp to blocks on each side. A smallhydraulic actuator on each side slides the lockingmechanism forward and aft (Fig 1-13). Two singlepush-pull hydraulic actuators, mounted on each side ofthe aft cargo compartment, lower and raise the ramp (Fig1-12). The ramp is hinged at the forward end to theaircraft structure. ADS links are installed at the aftcorners of the ramp and can be quickly released forlowering the ramp to the ground (Fig 1-13). These linksmust be connected for the sequence to continue so thepetal doors can open. During a normal open sequence(pressure door open and locked), the ramp is energizedclosed so that the ramp locks can be unlocked. The rampthen lowers to the ADS position. The ramp may belowered to the ground (after petal doors open) bydisconnecting the ADS links and lowering the rampusing the RAMP switch. During a normal closesequence, the ramp raises and remains energized untilthe pressure door is closed and locked. With the rampenergized up, the locking pins slide through the rampfittings and the blocks to lock the ramp. A ramp manualsafety pin can be installed at each side to provide apositive locked indication.

Petal DoorsTwo petal doors in the underside of the aft fuselagecomplete the cargo door system (Fig 1-8). When open,unobstructed access to the cargo compartment isprovided. When closed, the petal doors and rampprovide a complete aerodynamic closure over the aftfuselage opening. The petal doors are locked together bytwo lock assemblies mounted on the left petal door.When closed, the lock hooks engage fittings on the rightpetal door to lock them together. The lock assemblies areactuated by hydraulic actuators. A gearbox assembly,torque tubes, and jackscrew-type actuators drive thepetal doors opened or closed. Two selector valvescontrol the two lock actuators and hydraulic motor onthe gearbox.


1-10


1-11

Fig 1-8, Petal door components


1-12

Fig 1-9, Pressure door components

2

3

1. PIANO TYPE HINGE2. PRESSURE DOOR LOCK INDICATOR3. CAM JACK


1-13

Fig 1-10, Pressure door locks and aux latches

1

1. AUX LATCH2. PRESSURE DOOR3. DOOR LOCKS

2

3

1-5 INTERIOR DOORS SYSTEM

Three doors separate forward compartments from thecargo compartment. The flight station entrance doorprovides a closure between the flight station and cargocompartment. The lavatory door provides privacy from

the cargo compartment. The avionics access door ismounted behind the flight station ladder and providesaccess to the avionics bay.

1-6 DOOR WARNING SYSTEM

The door warning system provides a visual indicationwhen a door is not locked. The doors involved are thecrew door, both troop doors, pressure door, ramp, petaldoors, and stabilizer access door. Limit switches sensewhen any of these doors unlock, and turn on a NOTLOCKED light for the affected door. A DOOR OPEN

annunciator light will also come on. A bypass switch foreach door allows the indication to the annunciator lightto be disconnected, thus allowing monitoring of otherdoors. The door warning system also provides inputs tothe take-off warning system.


1-14

Fig 1-11, Pressure door actuator


1-15

Fig 1-12, Ramp actuators


1-16

Fig 1-13, Ramp components

1

2

1. ADS GUIDE LINKS2. RAMP LOCKS

SECTION II GROUND HANDLING,INSPECTION,

AND SERVICING

This section consists of information extracted from T.O. 1C-141B-2-00GE-00-1

After reading this section, you should be able to recall:

1. The description of:

• towing

• parking

• jacking

• inspections

• servicing

2. The location of:

• over the wing fuel caps

• single point refuel receptacles

• external power connection

• hydraulic cart connections

• engine oil service caps

Ground Handling, Inspection, and Servicing

2-1


2-2

2-1 GROUND HANDLING, INSPECTION, AND SERVICING

Flightline maintenance is an on going task that involvesalot of different activities. The following paragraphs willdiscuss:

• General Information

• Inspections

• Servicing

2-2 MAINTENANCE GENERAL INFORMATION

Safety! The first and foremost thought in your mind, asyou walk the ramp or through any maintenance shop,must always be SAFETY. Safety is more than followingtech data, more than the proper use of equipment. It’s amoral responsibility. You only need two things tomanage a good safety program: knowledge andenforcement. The knowledge comes from study andwork experience. Enforcement is personal and should bemandatory. Anytime you have that gut feeling thatsomething just is not right, go and investigate.

Every maintenance action on an Air Force base hasbehind it a T.O. or some written document telling youhow to operate and maintain the equipment. In thissection, you will find a general collection of proceduresthat you should be briefly familiar with now and becomean expert on in the near future. They are the commonday-to-day actions you will observe on the ramp. Thisbrief explanation will encourage you to spend some timein the T.O. library so you can enforce safety whileaccomplishing sound maintenance repair actions.

General Maintenance PracticesFor references, here are some general maintenancepractices that apply daily in the maintenance complex.

Whenever maintenance is required on any flight controlsystem, component, or tail area, the affected control

switch, lever, or wheel shall be tagged with a warning tagstat ing; “DO NOT OPERATE SYSTEM,MAINTENANCE IN PROGRESS”. Operation of thesystem with maintenance in progress or componentsremoved can result in injury to personnel or damage toequipment.

Disassemble equipment or components only as far asnecessary to replace defective components or to performnecessary maintenance, even though instructions formore extensive disassembly may be provided.

Tag for identification all disconnected electrical wires,connectors, hoses, and tubing. Refer to applicablewiring diagrams for correct installation of electricalleads. Be sure to cap off disconnected hoses and tubingto prevent entry of foreign materials into the system.

Tape exposed power leads to prevent short circuits.

Install new gaskets, O-rings, fittings, brackets, etc., inthe correct position as determined from the old unitbeing replaced.

Maintain adequate tube and hose clearances to preventchaffing, pinching, or pulling.

Safety wire nuts, bolts, fittings, and connectors inaccordance with Air Force procedures.

After completion of maintenance, check for signs ofabnormal operation, such as leakage, excessive heat, ornoise, and properly document the aircraft forms.

2-3 GROUND HANDLING

When you first hit the ramp in the morning, the aircraftare usually arranged for the days activity with hatchesand windows closed, pitot tube covers, and enginecovers installed. You’ll notice a fire extinguisherpositioned at each aircraft. A closer look at each aircraftwill reveal gear pins installed in the landing gear, withred steamers attached. “REMOVE BEFORE FLIGHT”streamers are attached to the pitot covers and engine

covers. Let’s cover some of the other areas involved inthe ground handling of aircraft.

TowingThe C-141B aircraft has the capability of being towed onpaved ramps and runways at a maximum grade of fourpercent. If it is necessary to tow the aircraft over roughsurfaces or snow and ice, the bridel method should beused. This consist of towing from the MLG gear with


2-3

two 3/4-inch cables (called a towing bridle). The aircraftmust be towed only when absolutely necessary to avoidundue stress on the MLGs.

ParkingIf there is a choice, head the aircraft into the wind. Afterthe aircraft is in position you may want to set the parkingbrakes. The MLG brakes may be mechanically set fortemporary parking. Never set the parking brakes whilethey are hot. Allow the brakes to cool for at least 15minutes before setting them. To set the parking brakes,depress the top of the rudder pedals, pull the parkingbrake handle full aft, release the pedals and then releasethe handle. When the brakes are set, the rudder pedalswill return approximately one inch. To release theparking brakes, depress the top of the rudder pedals andreturn the parking brake handle to the normal position.

JackingJacking a C-141B aircraft can be done three differentways: (1) full aircraft jack, (2) nose aircraft jack, and (3)axle jack. The type of jack depends on the maintenancerequired.

Fuel SystemFueling and defueling operations are normallyaccomplished through the single point receptacle located

on the outside of the right main landing gear wheel wellpod. The control panel is located at the flight engineer’sstation. Fuel distribution affects aircraft balance andcareful control must be maintained when fueling anddefueling the aircraft to maintain a safe center of gravitycondition.

Hydraulic SystemThere are four independent hydraulic systems thatsupply hydraulic fluid under pressure to thehydraulically operated components. They are referred toas No. 1, 2, 3, and 4 systems. There is a servicing centerlocated at each system reservoir. Normal groundoperation of the No. 3 hydraulic system is accomplishedthrough the use of two electrically driven pumps. No. 1and No. 2 are operated by engine driven pumps. No.4 isoperated by a hand pump. In the event this method is notavailable, there are external connections for No. 1 andNo. 2 systems located in the left and right wheel wellpods. There are two accumulators in No. 3 system andare used for APU starting and emergency brakeoperation.

Oxygen SystemThe aircraft is equipped with two independent highpressure liquid oxygen (LOX) systems. The systemsconsists of three LOX converters. One for the crewsystem which is located in the NLG wheel well. Two areused for the troop system which are located in the rightmain landing gear wheel well pod.

2-4 INSPECTIONS

Aircraft Inspections InformationPlans and Documentation tracks the time for schedulingaircraft inspection. They also schedule specialmaintenance actions such as time change items and timecompliance technical orders (TCTOs), to includeone-time inspections. The weekly and monthlymaintenance plans include the specific inspections ormaintenance actions required for each aircraft or specialpiece of equipment assigned during that period of time.The preflight inspection is generally the first inspectionprior to the aircraft assuming a mission. A preflight isusually is good for 72 hours as long as the aircraft isflown within 48 hours of the inspection. Next is the

thruflight, which would be performed between flights toensure the aircraft is ready for the next flight. At the endof the flying day a basic post flight (BPO) inspection isaccomplished. This inspection is usually done with apreflight included to inspect for servicing requirementsalong with preventive maintenance.

A home station check (HSC) is accomplished every 75days and is accomplished with the use of the BPOworkcards to include asterisk items. Also included is themajor and minor isochronical inspection system. This isaccomplished every 300 days, alternating between themajor and minor inspections.


2-4

2-5 SERVICING

Aircraft Servicing InformationAircraft are serviced with fuel, water, hydraulic fluid,engine oil, liquid oxygen (LOX), and gaseous nitrogen.There are hazards associated with all of these. If properT.O. procedures are not followed, loss of life or damageto equipment is very possible. Let’s look at the servicingpoints on the C-141B aircraft (Fig 2-1, 1 of 2 thur 2 of 2).

FuelThe aircraft can be refueled from the underground pitsystem or from trucks. A normal refuel requires threepeople qualified in the fuel system and safetyrequirements. The personnel will fill the positions of therefuel supervisor/fireguard, the panel operator, and therefueling nozzle operator/fireguard. The fuels systemoperator controls the pressure and flow from the cart ortruck to the aircraft. The supervisor is in charge of theoperation and stands at the fire bottle located at the noseof the aircraft. The panel operator will be in the flightengineer’s position to equally distribute the required fuelto the proper tanks. The nozzle operator stands close tothe SPR and monitors the SPR to ensure no leaks beginduring a fueling operation. Also, the aircraft can berefueled by using the over the wing method.

LOX ServicingLOX servicing can be another hazardous operation iftech data is not followed. Servicing consists of fillingtwo troop and one crew LOX converters.

HydraulicThe method of servicing the hydraulic systems is toservice their respective reservoir.

Nitrogen ServicingCompressed nitrogen is used to service tires,accumulators, and struts. Compressed air can be usedfor tires and struts; however, items serviced with airmust be purged and reserviced with nitrogen at the firstavailable opportunity.

Potable WaterWater is serviced through the tank filler valve on theforward right side of the fuselage. A spotter is to be usedwhen marshaling the truck towards the aircraft toprevent accidents.

Engine Oil ServiceEngine oil quantity is measured within 30 minutes ofengine shutdown. Oil samples are taken periodically andanalyzed to monitor engine wear. The joint oil analysisprogram (JOAP) should be conscientiously followed toreduce engine failures.

LubricationPerhaps one of the most important aspects of having anaircraft that is ready to meet any commitment is having itproperly lubricated. The reason we stress the word“properly lubricated” is that over lubrication can be asdestructive as no lubrication. Remember, lubricants thatare on exposed surfaces will attract dust that formsdestructive abrasive compounds. Some bearings thatrequire lubrication are sealed by the manufacturer. Noprovisions are made for further lubrication andlubrication should not be attempted. Other bearings arelubricated by hand packing, by gun through zerk andflush type lubrication fittings, and through oil reservoirs.Remember that low pressure grease guns should be usedto avoid “blowing” seals and bearing shielding.However, if a pressure gun is used, the technician shouldbe careful to limit the amount and quantity of greaseinjected.


2-5


2-6

Fig 2-1, (1 of 2) Servicing points

1. NO. 4 MAIN TANK FILLER CAP2. NO. 4 AUXILIARY TANK FILLER CAP3. RH EXTENDED RANGE TANK FILLER CAP4. NO. 3 AUXILIARY TANK FILLER CAP5. NO. 1 HYDRAULIC SYSTEM RESERVOIR6. NO. 3 HYDRAULIC SYSTEM ACCUMULATORS7. NO. 3 HYDRAULIC SYSTEM RESERVOIR8. NO. 2 HYDRAULIC SYSTEM RESERVOIR9. PETAL DOOR CENTRAL GEARBOX10. A-20 PORTABLE FIRE EXTINGUISHERS11. NO. 2 AUXILIARY TANK FILLER CAP12. LH EXTENDED RANGE TANK FILLER CAP13. NO. 1 AUXILIARY TANK FILLER CAP14. NO.1 MAIN TANK FILLER CAP15. FIRE EXTINGUISHER SERVICING INSPECTION

ACCESS16. CSD OIL AND THRUST REVERSER FILLER ACCESS17. ENGINE OIL FILLER ACCESS18. NO. 2 MAIN TANK FILLER CAP19. APU OIL FILLER ACCESS20. GALLEY WATER TANK21. GALLEY REFUSE CONTAINER

22. PORTABLE OXYGEN BOTTLE23. NO. 4 HYDRAULIC SYSTEM RESERVOIR24. CREW OXYGEN FILLER ACCESS25. BATTERY SUMP JAR26. FLIGHT ENGINEER’S FUEL MANAGEMENT PANEL27. LAVATORY WASH WATER TANK28. LAVATORY SERVICE CART CONNECTION29. ELECTRICAL SPARES BOX30. NO. 3 MAIN TANK FILLER CAP31. MAIN LANDING GEAR TIRES32. MAIN LANDING GEAR SHOCK STRUT33. MAIN LANDING GEAR BEAM POSITIONER34. NOSE LANDING GEAR TIRES35. NOSE LANDING GEAR STRUT36. CARGO COMPARTMENT OXYGEN FILLER ACCESS37. SINGLE POINT REFUELING ADAPTERS38. AIR CONDITIONING REFRIGERATOR COOLING

TURBINE39. APU FIRE EXTINGUISHER SERVICING DOOR40. STARTER OIL FILLER ACCESS


2-7

Fig, 2-1, (2 of 2) Servicing points


2-8

SECTION III EQUIPMENT ANDFURNISHINGS

This section consists of information extracted from T.O. 1C-141B-25GS-00-1.


1. The purpose of equipment and furnishings

2. The function of

• flight station equipment

• galley

• lavatory

• cargo compartment equipment

• emergency equipment

• aerial delivery system (ADS)

• special mission equipment


• flight station equipment

• galley

• lavatory

• cargo compartment equipment

• emergency equipment

• aerial delivery system (ADS)

• special mission equipment

Equipment and Furnishings

3-1


3-2

3-1 EQUIPMENT AND FURNISHINGS

Equipment and furnishings in the C-141B aircraft ensuremaximum accommodation for flight crew members.Facilities necessary to accomplish the assigned missionof the aircraft are also provided. Equipment andfurnishings for crew use include flight stationequipment, a galley, and a lavatory. Mission-oriented

features include cargo handling equipment, hardwareinstalled in the cargo compartment, and emergencyequipment for crew and personnel. Special equipment isincluded for aerial delivery of cargo. Provisions alsoexist for rigging the airplane as a troop or paratroopcarrier, and for air evacuation missions.

3-2 FLIGHT STATION EQUIPMENT

Flight station equipment consists of flight station seats,relief crew bunks and seat backs, consoles, and stowagecompartments. Worktables are provided for thenavigator and flight engineer. Other flight stationequipment included are flight station access facilities,crew movement aids, and various curtains for lightrestriction and privacy. Convenience facilities areprovided for communication equipment and personaleffects stowage. Provisions for checklists, chart holders,and technical orders are also included. Environmentalequipment includes sun visors and glare shields. Flightstation equipment supports primary flight crew membersin the performance of their duties. The equipmentconsists of crew member seats, worktables, consoles,and stowage compartments. Accommodations forauxiliary crew members include seats, bunks, andrestraint seat backs. General flight station equipmentincludes three types of seats and two bunks: upper andlower, for relief crew members. Side consoles at thepilot’s and copilot’s positions contain oxygen maskregulators and various control panels. An additionalconsole containing an oxygen regulator is installed at theauxiliary crew seat area. Map case and stowagecompartments are provided for the pilot, copilot, andnavigator. An additional stowage compartment islocated beneath the lower bunk. Additional equipmentand installations include an electrical spares box,scupper and seal installation, and technical orderstowage.

Flight Station SeatsThree types of seats are permanently installed in theflight station: one type for the pilot and copilot, a secondfor the navigator and flight engineer (Fig 3-1), and a thirdfor auxiliary crew members (Fig 3-2). The pilot’s andcopilot’s seats can be moved vertically, forward and aft,and laterally. The seats incorporate a removable andadjustable head rest, upward pivoting arm rests, and afour-position reclining backrest with verticaladjustment. The cushions are removable and velcro

upholstery is used. The navigator’s and flight engineer’sseats, mounted on stationary pedestals, provide forwardand aft movement of the seat on integral tracks, 360degree swivel capability, elevation control, recliningbackrest wi th vert ical adjustment, andremovable/adjustable headrest. The cushions areremovable and snap-on upholstery is used. An auxiliarycrew member seat is installed on each side of the flightstation entrance door. The seats are four-leggedaluminum seat frames which are secured to the floorstructure. A portable flight check seat is stowed underthe aft end of the navigator’s worktable. When in use, theseat is mounted aft of the center console. The tubularaluminum, three-legged fold up seat has a paddedbottom cushion and backrest and provision is made forsafety belt installation. Inertia reels are installed on thebacks on the pilot’s, copilot’s, navigator’s, and flightengineer’s seats. The inertia reels provide the restrainingforce on the shoulder harnesses installed on each of thefour seats. Safety belts are provided for the pilot’s,copilot’s, navigator’s, flight engineer’s, and auxiliarycrew member seats. Additional safety belt provisions aremade for lower bunk, upper bunk, and the flight checkseats.

Relief Crew BunksA double bunk is provided in the flight station for reliefcrew members (Fig 3-2). The bunk framework isconstructed of sheet metal and honeycomb and isequipped with two removable polyurethane foam padswith removable covers. The lower bunk curtain track ismounted on the lower pan of the upper bunk.

Side ConsolesThree side consoles are provided at the flight station.The fiberglasss consoles are located at the outboard sideof the pilot’s and copilot’s stations, and in the left rearcorner of the flight station (Fig 3-3).


3-3

Stowage CompartmentsFour stowage compartments are provided in the flightstation (Fig 3-3). The compartments include the pilot’s,copilot’s, and navigator’s stowage compartments and astowage area under the inboard portion of the lowerbunk.

WorktablesWorktables are provided at the navigator’s and flightengineer’s stations (Fig 3-4). The navigator’s worktableis constructed of sheet metal honeycomb and tubularaluminum and is secured to the flight station floor andfuselage structure. The flight engineer’s slide-mountedworktable is constructed of sheet metal honeycomb andaluminum and is an integral part of the flight engineer’sconsole. The table is fitted with a center pull out drawer.Stowage areas are provided for the flight engineer’s andnavigators oxygen mask, hose, and smoke mask.

Electrical Spares BoxAn electrical spares box, located in the upper right cornerof the navigator’s instrument panel, contains electricalspares for inflight use (Fig 3-4). Spares includecommonly used lights and fuses, and should be replacedwith new components as quickly as possible.

Technical Order (T.O.) File BookshelfA sheet metal bookshelf for T.O. file stowage is installedin the aft right corner of the flight station (Fig 3-2). Thebookshelf is divided into five vertical compartments andone horizontal compartment.

Miscellaneous Flight Station EquipmentMiscellaneous flight station equipment is provided forcrew movement safety, comfort, and convenience inperformance of assigned duties (Fig 3-5).

Flight Station Entrance LadderA three-step removable ladder is installed at the flightstation entrance for entering and exiting the flight station(Fig 3-5). The back of the ladder provides a closure forthe personnel opening into the electronic equipmentcompartment.

Hand Rails and Assist HandlesAssist handles are installed on either side of the flightstation entrance to facilitate entering and exiting theflight station (Fig 3-5). Hand rails are provided on right,forward, and left overheads as safety aids for personnelmovement in the flight station.

CurtainsCurtains are provided in the flight station for lightrestriction and personnel privacy (Fig 3-5). A smallcurtain is provided on the forward side of the flightstation door window for the flight crew privacy andrestriction of light from cargo compartment.

Hooks and HangersHooks and hangers are provided for storage or readyaccessibility to crew communications equipment and forconvenient stowage of crew clothing.

Checklist and Letdown Chart HoldersHolders for checklists and letdown charts are providedfor ready reference to flight information publications(Fig 3-5). The copilot’s spring-loaded, scroll-typechecklist holder is located over the copilot’s glare shield.The pilot’s and copilot’s letdown chart holders aremounted on the extreme left and right sides,respectively, of the main instrument panel glare shield.The holders are installed for ready reference to flightinformation publications and are manually movedbetween the stowed and viewing positions. (These arebeing phased out as supply diminishes).

Glare ShieldA three-piece glare shield assembly is installed over theinstrument panels to reduce glare on the panels. Theglare shield assembly is constructed of polyester glassfabric and painted black.

Sun VisorsThe pilot’s and copilot’s transparent green plastic sunvisors are installed above the windshields to reduceglare. When not required,the sun visors can be turned upout of the field of view.


3-4


3-5

Fig 3-1, Flight station seating


3-6

Fig 3-2, Relief crew seats and bunks

Fig 3-3, Flight deck stowage


3-7

Fig 3-4, Flight station work areas


3-8

Fig 3-5, Flight station equipment

CupholdersHot cupholders are provided at various locations in theflight station. The pilot’s and copilot’s cupholders are ontheir respective side console shrouds, and thenavigator’s holder is located on the forward side of hisworktable.

Fuel Tank DipstickThe fuel tank dipstick is stowed on the left panel of theentrance to the flight station above the ladder(Fig 3-5).

3-3 GALLEY

The galley provides a place for food storage andpreparation, hot beverage preparation, drinking waterstorage, and waste disposal. The galley consists of anoven, refrigerator, hot cup receptacle, hot beveragecompartment, waste compartment, worktable, controls,and indicator lights.

General Galley ProvisionsThe galley provides a work counter with formica coverand anti-drip edges. Also included are a built-in covered

refuse opening, a removable waste container, storagedrawer for utensils, a built-in light assembly over thework counter, and a hot cup receptacle. Majorcomponents can be removed individually or the galleycan be removed as a complete unit.

Fresh Water TanksTwo removable 2-gallon tanks mounted on racks abovethe galley provide drinking water for the crew and waterfor the hot beverage assembly when installed.


3-9

Fig 3-6, Galley and components

RefrigeratorThe refrigerator, located on the right side of the galleybeneath the work counter, provides storage for frozenmeals (Fig 3-6).

OvenThe electric warming oven is installed above the workcounter on the right center side of the galley (Fig 3-6).

Hot CupThe hot cup is a portable plug-in assembly used to heatfood and beverages in limited quantities (Fig 3-6).

Hot Beverage UnitProvisions for a hot beverage unit, to the left of thewarming oven, include water connection, power

receptacle, HOT BEVERAGE power switch, andindicator light (Fig 3-6).

Galley Control PanelPrimary control of the crew galley is through the galleypower switch located above the hot beverage chamber(Fig 3-6). A MAIN POWER indicator light is located tothe right of the power switch. Circuit protection of thegalley is provided by seven circuit breakers to the left ofthe hot beverage chamber. A GALLEY LIGHT switch islocated beneath the hot beverage chamber on the leftside.

Waste ContainerThe portable waste container, located to the left of therefrigerator and beneath the work counter, providesgalley refuse storage and water draining (Fig 3-6). Aplastic bag liner inserted in the container facilitatesservicing.

3-4 LAVATORY

A lavatory is provided for flight crew use. The lavatoryconsists of a toilet, wash basin, wash water supply, andwaste water storage. Other equipment includes a mirror,lighting, paper supplies and dispensers, and waste paperdisposal facilities. The crew lavatory is located on theairplane right side under the flight station floor. Thelavatory is accessible from the cargo compartment.Provision is made for a portable oxygen bottle.

ToiletThe flush-type toilet is attached to the aircraft structureon the right side of the lavatory (Fig 3-7). The flushingcontrol button is located on the sidewall over the toilet.

Wash WaterA 5-gallon fiberglass tank, installed in a compartmentabove the utility cabinet, stores wash water for the basin(Fig 3-7). A capped filler opening is provided in the tanktop for filling the tank when a service truck is notavailable.

Waste Water TankA 5.5 gallon fiberglass water tank is installed under thewash basin to hold waste water from the basin (Fig 3-7).

Waste Paper ContainerA waste paper container is provided in the lavatoryoutboard of the toilet.

Servicing ProvisionsThe drain control unit, installed near the toilet tank,provides for toilet tank drainage (Fig 3-7). The controlhouses a cap, petcock, and open/close mechanism. Thedrain control is released by turning a push-pull handleone quarter turn counterclockwise and pulling down tothe stop.

Miscellaneous Lavatory EquipmentA light fixture is provided above the adjustable mirrorwhich is installed above the wash basin in the lavatoryenclosure (Fig 3-7). A paper towel dispenser is installedabove the waste paper container on the inside of theservice access door. Towels are dispensed through a slotopening in the door front. The service door is secured inthe closed position by a latch. An assist handle isprovided on the front of the toilet tray beneath the mirror.A toilet tissue retainer and roller is provided below thewaste water tank. An electrical shaver outlet is installedon the panel above the toilet flush button. The electricalshaver circuit includes a power supply and converter inthe under deck equipment rack, which provides 115VAC, 60 Hertz (Hz) power to the outlet.


3-10

3-5 CARGO COMPARTMENT EQUIPMENT

The cargo compartment facilitates moving and securingcargo, and aerial delivery support equipment. Tiedownfittings, receptacles, and restraint equipment areincluded. Roller conveyers aid in cargo handling.

Walkways for personnel movement have special,non-skid material to ensure crew footing. Winches areprovided for cargo loading and other applications. Twoauxiliary ground loading ramps 8-1/2 ft. long by 22 in.


3-11

Fig 3-7, Lavatory components

wide are stowed on the sides above the cargo ramp.Three boxes for 10,000-pound tiedown chain stowageare installed on the right side of the fuselage at fuselagestations 530, 570, and 620. A fourth box on the right sideat fuselage station 710 provides stowage for 25,000-

pound tie-down chains. Two boxes are installed on theleft side at fuselage stations 590 and 620. These boxesprovide stowage for 25,000-pound tiedown chains. Theforward box also holds 10,000-pound tiedown fittingsand the parachute extraction line fittings.

3-6 EMERGENCY EQUIPMENT

Emergency equipment includes life rafts and attachedsurvival kits, escape ladders, escape ropes, crash axes,first-aid kits, fire extinguishers, life vests and anemergency locator transmitter (ELT). Six 20-man liferafts and attached survival kits are carried on the aircraft:four internally in the cargo compartment and twoexternally in covered wells in the right and leftwing-to-fuselage farings. Additional emergencyequipment is as follows:

Escape LaddersAn escape ladder is installed near each of the threeemergency escape hatches in the top of the cargocompartment. The ladders are constructed of knurledtubing and nylon rope and have nine rungsapproximately 12 inches apart.

Escape RopesAn escape rope is installed near each of the escapehatches in the cargo compartment. An escape rope is alsolocated near the escape hatch in the flight station.

Life VestsStowage pouches for life vests are provided on the backsof the pilot’s, copilot’s, auxiliary crew seats, and behindthe bunk seats.

First-Aid KitsTwo first-aid kits are installed on the bulkhead behindthe outboard auxiliary seat in the flight station. Four kitsare installed in the cargo compartment; two are locatedjust aft of the crew entrance door and the other two areinstalled just aft of the two troop doors. Provision ismade for installation of 20 additional kits in the cargocompartment. Five kits can be installed immediately aftof each of the four side emergency exit doors.

Fire ExtinguishersSix halon 1211 fire extinguishers are carried on theaircraft: one is installed in the flight station, two areinstalled just aft of the crew entrance door, one is locatedmidway in the cargo compartment on the right side, andtwo are installed forward of the left troop door.

Crash AxA crash ax is stowed on the bulkhead behind the inboardauxiliary seat.

Cargo Compartment Warning HornA warning horn is installed overhead in the cargocompartment at fuselage station 1058. The horn ismanually controlled in emergency conditions by theBAIL OUT ALARM switch on the pilot’s overheadpanel. The horn is automatically activated by a flow ofoxygen in the troop oxygen system. The warning horn isalso connected to the APU fire warning circuit to warn ofan APU fire on the ground.

3-7 AERIAL DELIVERY SYSTEM (ADS)

The aerial delivery system supports cargo air drop. Anextraction parachute release mechanism releases theparachute electromechanically or manually (Fig 3-8).Electromechanical release is accomplished byenergizing a release solenoid which drives a release pinagainst a lever on the parachute holder to release theparachute. Manual release is effected by pulling a D

handle located below the crew door interphone andpublic address (PA) panel. The handle is connected by acable to a lever which is forced down against the releasepin flange. A parachute winch and parachute releasependulum facilitate parachute handling and release,respectively. Bailout alarm and troop jump lights arealso installed.


3-12


3-13

Fig 3-8, Aerial delivery system control panel

3-8 SPECIAL MISSIONS EQUIPMENT

The aircraft can be configured for special missions byinstalling special kits. These kits include the following:oxygen supply and distribution kit, life raft liner kit,stanchion kit, litter installation kit, troop transport kit,and paratroop kit. Paratroop spoiler doors, jumpplatforms, and anchor cables are also included. Bothintegral and removable cable supports as well ashandling winches are provided.

Oxygen Distribution and Supply KitWhen the aircraft is used as a troop carrier or airevacuation, oxygen must be available for all personnel inthe event of loss of cabin pressure. Tubing, connectors,fittings, and other oxygen equipment are permanentlyinstalled in the aircraft. A removable oxygen anddistribution supply kit is installed. The kit uses two75-liter converters which provide 460 manhours ofoxygen at cruise altitude. The kit also provides alloxygen masks and associated equipment required for afull complement of troops.

Life Raft Liner KitThe life raft liner kit provides all hardware needed toinstall six 20-man life rafts and survival kits in theaircraft. The kit consists of wing compartment liners,containers for the two aft fuselage upperdeck life rafts,and holding harnesses and cradles for the emergency exitlife rafts and survival kits.

Stanchion KitStanchions must be installed before troop seats or littersare installed in the aircraft. A stanchion kit contains thecenter and sidewall stanchions and attachingcomponents required for installing troop seats or litters.Sidewall stanchions are used for litter installation only.Center stanchions may be used for troop seats or litters.

Comfort PalletProvisions for a comfort pallet are installed in theforward end of the cargo compartment (Fig 3-9). Thecomfort pallet is used to increase meal and lavatorycapacity when the aircraft is equipped for trooptransport. The comfort pallet is mounted on a cargopallet for ease of installation and removal.

Litter ConfigurationThe litter provision kit provides for the installation of103 litters. The centerline and sidewall stanchion kits arerequired to install the litters. Additional seats forattendants can be installed as required.

Troop SeatsTwo types of seats are installed for personnel transport:rigid aft facing seats (Fig 3-9) and canvas sidewall andcenterline seats.

Paratroop Seat ConfigurationProvisions for equipping the aircraft for a paratroop dropmission are included in a paratroop kit. The paratroop kitincludes canvas seats, static line anchor cable kit,electrically actuated spoiler doors, jump platforms, and aroller bar to assist in paratrooper retrieval.

Paratroop Spoiler DoorsRight and left spoiler doors for use during paratroopdrops are supplied in the paratroop kit. Each door is anair deflecting surface approximately 2 by 4 ft.

Anchor CablesTwo short anchor cables are in the paratroop kit and canbe used on all aircraft. Two long anchor cables arestowed on reels and mount on the right side of the cargocompartment between fuselage stations 538 and 558.

Portable Urinal (Honey Bucket)A portable urinal housed in a light weight plasticcontainer is in the paratroop kit. The unit consists of twourinals, drain hose, and drain tanks. The capacity of eachtank is 6.5 gallons. A valve in each hose can be used toshutoff the bowl drains while the drain tanks are beingemptied. The drain tanks and a hinged front panelcovering the drain tanks are held in place withwebstraps. A handlebar across the top of the containerprovides a hand hold.


3-14


3-15

Fig 3-9, Special mission kits

1. COMFORT PALLET2. AFT FACING SEATS3. SEAT MOUNTING TRACKS


3-16

SECTION IV LIGHTING SYSTEM

This section consists of information extracted from T.O. 1C-141B-2-33GS-001.


1. The purpose of the lighting system

2. The function of

• flight station lighting system

• cargo and service compartment lighting system

• exterior lighting system

• emergency lighting system


• flight station lighting system

• cargo and service compartment lighting system

• exterior lighting system

• emergency lighting system

Aircraft Lighting

4-1


4-2

4-1 LIGHTING SYSTEM

The C-141B aircraft lighting system consists of theinterior lighting (flight station and cargo compartment),exterior lighting, and 11 emergency exit lights. Theinterior lighting provides general illumination of theflight station, cargo compartment, supplementaryinstrument lighting, and lighting aft of the aircraft during

cargo loading. Exterior lighting provides light forlanding, taxiing, takeoff, wing inspection, formationflying, recognition purposes, and aerial refueling. Anemergency exit light is near each emergency exit and atthe troop and crew entrance doors.

4-2 FLIGHT STATION LIGHTING SYSTEM

The flight station lighting system provides generalillumination of the flight station and supplementarylighting of instruments and control panels (Fig 4-1).Overhead dome lights, instrument lights, instrumentpanel lights, utility lights, and reading lights comprise

the flight station lighting. A master caution warningsystem, consisting of a combined light and reset switchon the pilot’s and copilot’s instrument panel, centerconsole annunciator lights, and a control is included inthe flight station.

4-3 CARGO AND SERVICE COMPARTMENT LIGHTING SYSTEM

The cargo and service compartment lighting systemcovers all lighting for the cargo compartment andhousing of various components and accessories. Generallighting throughout the cargo compartment is providedby overhead dome lights controlled by switches on theforward crew interphone and PA panel, the ramp anddoor control panel, or the flight station door switchpanel. Service lighting is available in the lavatory, underflight deck rack area, aft crawl way and vertical stabilizertunnel, nose and main wheel wells, cargo loading area,

and paratroop jump platforms. Cargo and servicecompartment lighting is achieved by cargo compartmentdome and overhead lights, crew rest platform lights, alavatory light, a flight station access light, ramp loadinglights, nose and main wheel well lights, under flight deckrack lights, vertical stabilizer lights, paratroop jumpplatform lights, and their respective switches.

Aircraft Lighting

4-3

1. LOWER BUNK READING LIGHT2. UPPER BUNK READING LIGHT3. OUTBOARD SEAT READING LIGHT4. PILOT’S UTILITY LIGHT5. NAVIGATOR’S UTILITY LIGHT6. NAVIGATOR’S WORKTABLE LIGHT7. DOME LIGHT

4

5

6

72

3

Fig 4-1, Flight station lighting

Cargo Compartment Dome LightsThey consist of 63 white lights and 22 red lights in thecargo compartment ceiling (Fig 4-2). The dome lightsare normally controlled by the CARGO COMP LIGHTScontrol and dimming switches on the forward crewinterphone and PA panel.

Cargo Compartment Overhead LightsThirty eight white, non-dimmable, dome-type lights areinstalled in the cargo compartment overhead (Fig 4-2).


4-4

Fig 4-2, Cargo compartment overhead lights

Crew Rest PlatformTwo crew rest platform lights are located on both sides ofthe upper fuselage (Fig 4-3).

Lavatory LightThe lavatory light switch is on the left side of the cantedceiling panel inside the lavatory door. The switchcontrols the lavatory light on the bulkhead above thesink and the light above the latrine door.

Aircraft Lighting

4-5

Fig 4-3, Crew rest platform lights

1. TYPICAL CREW REST PLATFORM LIGHT2. LADDER LIGHTS SWITCH3. LATRINE DOOR LIGHT

Fig 4-4, Lavatory and avionics compartment lights

1. AVIONICSCOMPTARTMENTLIGHT

2. NOSE WHEELLIGHT SWITCH

3. LAVATORY LIGHT

Flight Station Access LightOne flight station access light is to the right of the flightstation access door. The light is controlled by theLADDER LIGHT switch to the left of the flight stationentrance door.

Cargo Ramp Loading LightsTwo adjustable cargo ramp loading lights are near theforward end of the aft cargo deck . These lights aremounted at the left and right outboard edges of the aftcargo deck. The cargo ramp loading lights are controlledby the RAMP LOADING LIGHTS switch on the doorand ramp control panel (Fig 4-5).


4-6

Fig 4-5, Cargo compartment ramp loading lights

Nose and Main Wheel Well LightsOne nose wheel well light is installed on the inside of thenose wheel well for illumination to inspect the noselanding gear down lock. The nose wheel well light iscontrolled by the WHEEL WELL LIGHT switch besidethe nose wheel well observation window (Fig 4-4). Theleft and right main wheel wells each contain a mainwheel well light. A separate control switch for each lightis adjacent to the left and right main landing gearobservation windows at fuselage station 1030 .

Under Flight Deck Rack LightsTwo lights are installed in the under flight deck area forillumination of the avionics and electrical racks. Acontrol switch is adjacent to each light (Fig 4-4).

Vertical Stabilizer LightsTwo inspection lights are located in the verticalstabilizer tunnel and four lights are located in the aftcrawl space.

Platform LightsOne red paratroop platform light is installedimmediately aft of each aft troop door. The left paratroopplatform light is controlled by the PLATFORMLIGHTS switch on the left jumpmaster/loadmasterpanel; the right paratroop platform light is controlled bythe PLATFORM LIGHTS switch on the rightjumpmaster/loadmaster panel.

4-4 EXTERIOR LIGHTING SYSTEM

The exterior lighting system provides light for landing,taxing, wing inspection, takeoff, formation flying, aerialrefueling, and recognition purposes. Control switchesfor exterior lighting are located on the pilots’ overheadpanel, except aerial refueling light controls are located atthe flight engineer’s station.

Landing LightsA sealed beam landing and terrain clearance light(landing light) is mounted on the bottom of each wingbetween the engine nacelles (Fig 4-6).

Formation LightsNine lunar white lights on the top of the aircraft and acontrol switch on the overhead panel comprise theformation lights system. Three lights are on each of theouter wing panels and three lights are on the fuselage(Fig 4-8).

Navigation LightsThe navigation light system consists of a red light on theleft wing tip, a green light on the right wing tip, white taillights, and a control switch on the pilots’ overhead panel(Fig 4-7).

Anti-Collision LightsThe aircraft is equipped with three anti-collision lights,one rotating beacon on the upper surface of the

horizontal stabilizer, one strobe light assembly on thetop of the fuselage in line with the wing, and one strobelight assembly on the lower fuselage at approximatelythe same line (Fig 4-7).

Taxi LightsThe taxi light system uses four light fixtures mounted onthe main landing gear doors (Fig 4-6). In addition, theleading edge lights are used in conjunction with taxilights.

Wing Leading Edge LightsA light is installed on each side of the forward fuselage ina position which will illuminate the engine nacelles andthe immediate leading edge area of each wing duringtaxi and normal flight operations not involving aerialrefueling (Fig 4-6).

Aerial Refueling Fairing LightsThree fairing lights are located on the forward fuselagejust ahead of the fairing. One light is located directly infront of the slipway door, and one on each side of thefuselage directed at the UARRSI fairing (Fig 4-8).

Aerial Refueling Slipway LightsTwelve slipway lights illuminate the slipway door area.Six are located on either side of the slipway (Fig 4-8).

Aircraft Lighting

4-7


4-8

Fig 4-6, Exterior lighting

FORMATION

LEADING EDGE

ANTI-COLLISION

NAVIGATION

Aircraft Lighting

4-9

2

Fig 4-7, Exterior lighting

1. WING NAVIGATION LIGHT2. FORMATION LIGHT3. UPPER ANTI-COLLISION STROBE LIGHT4. UPPER ANTI-COLLISION LIGHT (ROTATING

BEACON)5. LOWER ANTI-COLLISION STROBE LIGHT

1

4-5 EMERGENCY LIGHTING SYSTEM

The emergency lighting system consists of 11emergency exit lights located at each emergency exit andat the crew and troop doors (Fig 4-9). Each emergencyexit light contains batteries, relays, and a three-positioncontrol switch marked ARMED, OFF, and ON. For

normal operation, each control switch is placed toARMED. The lamp of an emergency exit light may beturned on at any time by setting the control switch of thatemergency exit light to ON. Stowing the handle causesthe light to go out.


4-10

Fig 4-8, A/R and formation lights

1. FORMATION LIGHTS2. SLIPWAY LIGHTS3. AERIAL REFUELING LIGHTS CONTROL

PANEL4. FAIRING LIGHTS

3

1 1

4 2

Aircraft Lighting

4-11

Fig 4-9, Emergency exit light


4-12

SECTION V FIRE PROTECTIONSYSTEM

This section consist of information extracted from T.O. 1C-141B-2-26GS-00-1.


1. The purpose of the fire protection system

2. The function of

• detection and warning systems

• fire extinguishing systems


• detection and warning systems

• fire extinguishing systems

Fire Protection

5-1


5-2

5-1 FIRE PROTECTION SYSTEM

The C-141B fire protection system consists of detectionand warning systems, fire isolation systems, and fireextinguishing systems. Fire warning is provided for eachof the four engine nacelles and the APU compartment.Overheat warning is provided for each of the four pylons

and the bleed air system. Smoke detectors are installed inthe cargo compartment. Fire extinguishing systems areprovided for the engine nacelles and the APUcompartment.

Fire Protection

5-3

Fig 5-1, APU sensing element

1. APU FIRE SENSING ELEMENT


5-4

Fig 5-2, Cowl door and pylon detectors

5-2 DETECTION AND WARNING SYSTEMS

The fire detection system senses the presence of a fire ineach engine nacelle and the APU compartment (Fig 5-1and 5-2). The overheat system senses the presence of anoverheat condition in each pylon. A fire condition isindicated by a steady light in the fire emergency controlhandle and a steady light in the master fire light (Fig 5-3and 5-4). Pylon overheat is indicated by alternateflashing of the master FIRE warning lights and thecontrol handle light for that engine. The master FIREwarning lights are located on the pilot’s and copilot’sinstrument panels. Each light is common to all fourengine and pylon detection systems (Fig 5-2). A fireemergency control handle for each engine is on the fireemergency shutdown panel. It is located above the maininstrument panel. Control handles for the APU arelocated at the flight engineer’s station and in the cargocompartment just aft of the crew door. An APU FIRElight is on the annunciator panel. Test switches forENGINE FIRE TEST and PYLON OVERHEAT TESTare located on the pilot’s pedestal. The APU FIREWARN TEST switch is on the flight engineer’sinstrument panel. An audible fire warning signalprovides a signal through the interphone system. Thesignal is also produced by the loudspeaker in the flightstation. This warning is common to all four engines and

the APU detector systems. The APU audible alarmsystem includes the cargo compartment warning horn. Itsounds if an APU fire occurs on the ground. The audiblewarning system consists of a fire warning generator,audible signal silencing relay, audible fire warningsilencing switch, and an amplifier for each crew station,except the navigator’s. These amplifiers are also used inthe maximum speed warning system. All of the audiblesignals are switched off by using the AUDIBLE FIREALARM SILENCE switch. It is located on the enginefire emergency control panel. A smoke detection systemprovides CARGO SMOKE warning lights on the flightengineer’s panel and the annunciator panel. Onedetector is mounted under the flight deck. Five detectorsare mounted in varying parts of the cargo compartment.A test switch is located on the flight engineer’s panel.Independent left, right, and cargo floor bleed airoverheat detection systems are provided. They consist ofa continuous loop type sensor, a control box, warninglight, and a reset and test switch. When an overheatcondition is detected, the system automatically closesvalves to isolate the affected system. The warning lightalso turns on.

Fire Protection

5-5


5-6

Fig 5-3, APU fire handles

5-3 FIRE EXTINGUISHING SYSTEMS

The engine nacelles and APU compartment have fireextinguishing systems. Extinguishing agent isdischarged into the fire zone by actuating the agentdischarge switch. This occurs after the fire emergencycontrol handle has been pulled (Fig 5-4). The fireextinguishing agent used is halon. It is contained inbottles mounted in the aft end of the outboard pylons.The agent for the APU compartment is contained in abottle mounted in the left wheel well compartment . Theagent in the left outboard pylon is used to extinguishfires in the No. 1 or No. 2 engine. The agent in the rightoutboard pylon is used for the No. 3 or No. 4 engine. Thespherical containers used for nacelle fire extinguishing

agent are made of stainless steel (Fig 5-5). Eachcontainer has two discharge heads. Nitrogen is used topressurize the container to a pressure of 600 PSI (plus 25or minus 0) at 72oF. They are mounted aft of the pylonbulkheads in the pylon box structure. A large accesspanel is on the left side of the pylons for servicing thecontainers. Control of the engine nacelle fireextinguishing system is by four agent dischargeswitches. The four agent discharge switches are locatedbehind the four engine fire emergency control handles.When the control handles are in the normal position, theagent switches are covered by the control handles. Whena control handle is pulled, its agent discharge switch is

Fire Protection

5-7

Fig 5-4, Engine fire handles

exposed. The two selector switches determine whichbottle will be discharged into the engine nacelle. Eachswitch has two posit ions: ALTERNATE andNORMAL. The APU fire extinguisher is controlled byAPU fire emergency control handles and the APUAGENT DISCHARGE switches (Fig 5-3). These itemsare located on the flight engineer’s panel and in the cargocompartment . The discharge switch is next to the controlhandle. However, the discharge switch cannot fire the

squib unless the control handle has been pulled. If a fireoccurs in the APU, one of the control handles is pulled.This shuts down the APU. It also closes the APU doorand completes a circuit from the isolated DC bus to thedischarge switch. When the discharge switch is pressed,a circuit is completed to the squib on the bottle. There arealso six portable extinguishers installed in the flightstation and cargo compartment for use by the crew.


5-8

Fig 5-5, Engine fire discharge bottle

SECTION VI LANDING GEARSYSTEM



1. The purpose of the landing gear system

2. The function of

• main landing gear and doors system

• nose landing gear and doors system

• extension and retraction system

• main landing gear brakes and anti-skid system

• nose landing gear steering system

• landing gear position and warning system


• main landing gear and doors system

• nose landing gear and doors system

• extension and retraction system

• main landing gear brakes and anti-skid system

• nose landing gear steering system

• landing gear position and warning system

Landing Gear Systems

6-1


6-2

6-1 LANDING GEAR SYSTEM

The C-141 aircraft landing gear is a modified tricycledesign with a steerable two-wheel nose gear and twobogie type four-wheel main gear. All three landing gearsare fully retractable. Hydraulic actuators retract the gearup and forward. The wheel well doors are connected tothe landing gear by rods, torque tubes, and bellcrankscausing the doors to open as the gear extends and to close

as it retracts. The two main landing gears have discbrakes on each wheel. An anti-skid system is used toimprove braking performance. The nose landing gear issteered hydraulically. Normal hydraulic power forextension, retraction, braking, and steering is providedby the No. 2 hydraulic system.

6-2 MAIN LANDING GEAR AND DOORS SYSTEM

The main landing gear are installed in pods on the leftand right sides of the lower center fuselage. Each mainlanding gear includes a shock strut, drag braces,actuators, torque arms, an axle beam (commonly called abogie), wheel and tire assemblies, and brake assemblies.

Shock StrutThe shock strut is an air-oil cylinder and piston assemblythat absorbs shocks during taxiing, take-off, and landing.The shock strut pivots on a trunnion shaft which attachesthe main landing gear to the aircraft (Fig 6-1).

AxlesThe axles are mounted on a tubular, bogie-type axlebeam that pivots on the base of the shock strut (Fig 6-1).The axle beam positioner, a small air-oil cylinder, holdsthe axle beam perpendicular to the shock strut when thegear is extended. The axle beam is kept parallel to thefuselage by a hinged torque arm assembly. A leveler rodattached to the upper torque arm moves the axle beaminto a level position as the gear nears full retraction. Thegear is extended and retracted through the use of afolding upper drag brace assembly. A smaller downlockactuator locks the gear in the extended position bypreventing the upper drag brace assembly from folding.Three lower drag braces attach to the trunnion shaft andhelp brace the gear to the aircraft structure.

Wheels, Tires, And BrakesEach main gear has four identical wheel and tireassemblies, each of which rotates on two sets of taperedroller bearings (Fig 6-1). Anti-skid detectors aremounted in both ends of the two axles. The multi- discbrake assembly on each wheel is braced to the shockstrut by a brake torque link. The torque links prevent theaxle beam from pitching forward when the brakes areapplied.

DoorsEach main landing gear wheel well has one upper andtwo lower doors. The shock strut is connected throughballjoint and bellcrank assemblies to a torque tuberunning the length of the wheel well. When the gear isextended or retracted, movement of the shock strut istransmitted through the torque tube to pushrods thatopen and close the lower doors (Fig 6-2). A bellcrank onthe torque tube hooks into the main landing gear uplockassembly to lock the gear in the fully retracted position.An additional door uplock latch assembly (commonlycalled a flapper door) ensures that the lower doorsremain fully closed in flight. When the gear is extended,a pushrod connected to the shock strut opens the oleo(upper) door to allow clearance between the shock strutand the top of the wheel well.


6-3


6-4

Fig 6-1, Main landing gear

13. OUTBOARD LOWERDRAG BRACE

14. BRAKE ASSEMBLY15. AFT AXLE16. WHEEL ASSEMBLY17. BRAKE TORQUE LINK18. AXLE BEAM (BOGIE)19. TIRE20. ANTISKID DETECTOR21. UPPER TORQUE ARM22. LOWER TORQUE ARM23. TORQUE TUBE24. UPLOCK ASSEMBLY

1. SHOCK STRUT2. LOWER DRAG BRACE3. OLEO DOOR PUSHROD4. FORWARD UPPER DRAG BRACE5. TRUNNION SHAFT6. INBOARD LOWER DRAG BRACE7. DOWNLOCK ACTUATOR8. AFT UPPER DRAG BRACE9. MLG ACTUATOR10. ATTACH LINE LINK ASSEMBLY11. TUBE BRACE12. YOKE

25. UPLOCK BELLCRANK26. BELLCRANK DRIVE

BALL JOINT ASSEMBLY27. BELLCRANK ASSEMBLY

6-3 NOSE LANDING GEAR AND DOORS SYSTEM

The nose landing gear uses an air-oil shock strut tocushion the jolts of taxiing, take-off, and landing. Thehorizontal shock strut trunnion connects the nose gear tothe aircraft structure and is mounted in bearings on eachend. The bearings allow the shock strut trunnion to berotated by a push-pull hydraulic actuator to extend andretract the gear. A built-in up/down lock mechanism islocked and unlocked by a small up/down lock actuatormounted near the top of the shock strut (Fig 6-3). Thenose gear can be steered with the steering wheel or therudder pedals. A large steering actuator is mounted justforward of the shock strut trunnion. When the gear is

retracted, spin brakes rub against the tires to stop thewheels from spinning.

DoorsThe nose landing gear wheel well is enclosed by threedoors (Fig 6-4). The two forward doors open down andoutward as the gear is lowered, then close again when itis fully extended. The aft door opens by moving downand aft along the fuselage. An aerodynamic baffle doorkeeps the aft door from catching the wind.


6-5

Fig 6-2, Main landing gear doors

1. OLEO DOOR2. OUTBOARD DOOR3. FORWARD PUSHROD (OUTBOARD DOOR)4. DOOR UPLOCK LATCH ASSEMBLY (FLAPPER)5. FORWARD PUSHROD (INBOARD DOOR)6. INBOARD DOOR


6-6

Fig 6-3, NLG support and actuating components

13. DISCONNECT PINS14. TORQUE ARM ASSEMBLIES15. NLG ACTUATOR16. SPHERICAL17. SHOCK STRUT TRUNNION18. SHOCK STRUT

1. RIGHT NLG CRANK2. UP/DOWN LOCK ACTUATOR3. LEFT NLG CRANK4. UP/DOWN LOCK PUSHROD5. UP/DOWN LOCK EMERGENCY VALVE6. UP/DOWN LOCK EMERGENCY RELEASE

LEVER7. FILLER VALVE8. DRAG BRACE ASSEMBLY9. SPHERICAL

10. BEARING11. MLG TRUNNION12. BEARING


6-7

Fig 6-4, NLG doors and drive mechanism

1. LEFT FORWARD DOOR2. RIGHT FORWARD DOOR3. FORWARD PUSHROD4. BELLCRANK5. SUPPORT TUBES6. PUSHROD7. TRUNNION BELLCRANK8. SHOCK STRUT TRUNNION9. TRUNNION BELLCRANK10. ATTACH FITTING11. AFT PUSHROD12. FORWARD ARM13. ATTACH FITTING14. AFT ARM15. AFT DOOR

16. BUMPER17. BAFFLE DOOR18. BUMPER19. AFT ARM20. ATTACH FITTING21. AFT PUSHROD22. TRUNNION BELLCRANK23. ATTACH FITTING24. TRUNNION BELLCRANK25. PUSHROD26. SUPPORT27. FORWARD PUSHROD28. BELLCRANK

6-4 EXTENSION AND RETRACTION SYSTEM

The main landing gear and nose landing gear areextended or retracted simultaneously. Normal extensionand retraction is controlled electrically and powered bythe No. 2 hydraulic system. Movement of the landinggear control handle causes selector valves to routehydraulic pressure to the appropriate actuators and movethe gear. If hydraulic pressure is lost or electricalcontrols fail, emergency extension systems can be usedto lower the gear manually.

Normal Extension and RetractionThe nose landing gear selector valve, main landing gearselector valve, main landing gear downlock selectorvalve, and main landing gear doors uplock selector valvedirect hydraulic pressure for landing gear operation (Fig6-5). All four valves are solenoid actuated with manualoverrides provided in case of electrical failure. The mainlanding gear selector valve, located on the left side of thecargo compartment, controls the flow of hydraulicpressure to the main landing gear actuators. Flowregulators are used to control the speed of extension andretraction. The downlock actuator locks the drag linkassembly over center when the gear is fully extended.The nose landing gear selector valve is located in theelectronics compartment below the flight station near thecrew entry ladder. It directs hydraulic pressure to thenose landing gear actuator, up/down lock actuator,steering actuator, and the normal brake selector valve.

Emergency ExtensionThe main landing gear emergency extension system isused to release and lock down the main gear when ahydraulic failure will not allow for normal extension.Each main gear has its own emergency extension systemconsisting of three handles for each main landing gear.They are located on both sides of the cargo compartmentjust aft of the hydraulic servicing center. The step onehandle is pulled to release the flapper door uplock, andstep two handle is pulled to release the main gear uplock(Fig 6-7). When the gear has fallen into position, the stepthree handle and bellcrank mechanism is used to lock thedrag link assembly over center (Fig 6-6).

The manually operated number four hydraulic systemprovides emergency extension for the nose landing gear.The system is located below the flight station on the rightside just forward of the lavatory compartment. Itconsists of a reservoir, an emergency selector valve, ahand pump, a relief valve, and associated plumbing. Ifthe No. 2 hydraulic system should fail, the hand pumpcan be used to lower the gear. If the hand pump shouldfail, an access panel in the left avionics bay below theflight station can be removed to gain access to theemergency release lever located on the NLG strut.Pulling the lever allows the nose landing gear to free fall(Fig 6-3).

6-5 MAIN LANDING GEAR BRAKES AND ANTI-SKID SYSTEM

The main landing gear brakes are operated hydraulicallyusing No. 2 system pressure for normal operation andNo. 3 system pressure for emergency braking. Ananti-skid system can be used with the normal brakesystem to prevent locked wheels and skidding. The brakeand anti-skid panel, located on the right side of the centerinstrument panel, has an anti-skid switch and a brakeselector switch with pressure gages for normal andemergency brake systems. Parking brakes are set with aT-handle on the left side of the pilot’s instrument panel.

BrakesEach main landing gear wheel has its own multi-discbrake assembly (Fig 6-8). The main landing gear brakesystem has two dual pilot metering valves, one for eachgear. The dual pilot metering valves transformmovement of the brake pedal linkage into meteredhydraulic pressure. That pressure is supplied from the

No. 2 hydraulic system. When the brake selector switchis placed to the EMER position, the normal brakeselector valve shuts off No. 2 pressure and theemergency brake selector valve sends No. 3 hydraulicsystem . During normal braking, the pilot pressure fromeach metering valve is routed to a pair of anti-skidcontrol valves. For emergency braking, the anti-skidcontrol valves are bypassed. No. 3 system pilot pressureis directed to a main brake metering valve that sends No.3 system pressure to the brake assemblies.

Anti-SkidThe main landing gear normal brakes are equipped withan anti-skid system to prevent the dangerous loss ofbraking and control that occurs during a skid. Thesystem works by sensing a skid in its earliest stages.Each main gear wheel rotates the shaft of an anti-skiddetector (Fig 6-1). The anti-skid detector generates an


6-8

electrical signal based on that rotation and sends thatsignal to the anti-skid control box located in the leftcargo compartment. The anti-skid system operates untilthe aircraft has slowed below 15 knots. A fail-safefeature prevents an anti-skid malfunction from causing aprolonged loss of braking action. The anti-skid system

also works to ensure that all brakes are fully releaseduntil after the aircraft touches down. When the landinggear is extended during approach, the BRAKES RELlight on the brake and anti-skid panel comes on toindicate that the brakes are released.


6-9

Fig 6-5, Hydraulic selector valve

1. MLG DOOR UPLOCK SELECTOR VALVE2. MLG DOWNLOCK SELECTOR VALVE3. MLG SELECTOR VALVE


6-10

Fig 6-6, Main landing gear emergency extension components

1. STEP 1 HANDLE ( DOOR UPLOCK LATCH EMERGENCY RELEASE)2. STEP 2 HANDLE (UPLOCK EMERGENCY RELEASE)3. STEP 3 (EMERGENCY DOWNLOCK) HANDLE


6-11

Fig 6-7, Emergency extension components

1. RELEASE ARM2. UPLOCK EMERGENCY RELEASE CABLE3. MLG GROUND SAFETY PIN4. SHAFT5. EMERGENCY DOWNLOCK FITTING6. BELLCRANK7. RELEASE ARM8. DOOR UPLOCK LATCH EMERGENCY CABLE


6-12

Fig 6-8, Brake assembly

1. BLEED VALVE2. BRAKE ASSEMBLY3. DISK4. PISTON (11 PLACES)5. RELEASE SPRING HOUSING

6-6 NOSE LANDING GEAR STEERING SYSTEM

The aircraft can be steered by either moving the rudderpedals or rotating the steering wheel located to the left ofthe pilot’s instrument panel. The motion is relayedthrough a series of chains, pulleys, and cables to asteering control valve mounted over the steeringactuator. The steering control valve responds to cablemovement by sending No. 2 system hydraulic pressureto the appropriate end of the steering actuator. Theactuator contains opposing pistons connected by a rackgear. The rack gear meshes with a sector gear that rotateson the shock strut cylinder. Rotation of the sector gearsteers the wheels by means of a torque arm assembly thatconnects to the nose gear axle. The rudder pedals areused to steer the aircraft at higher speeds and can onlyturn the wheels up to eight degrees left or right. Thesteering wheel is used to steer the aircraft while taxiing atlower speeds. It allows the wheels to be turned up to 60degrees in either direction. The steering wheel is

mounted on an upper steering column. A gearboxconnects the upper and lower steering columns andmakes it easier to turn the steering wheel. The gearboxalso contains a position indicator that shows how far leftor right the nose wheels are turned. An interconnectassembly links the rudder pedal steering and steeringwheel components. It contains a cam and sprocketcombination that is connected with a drive chain to thelower steering column. The interconnect assemblyallows the nose gear to be steered eight degrees in eitherdirection using the rudder pedals. The interconnectassembly is located below the flight station floor to theleft of the pilot’s control column. When the aircraft takesoff, the shock strut extends and the wheels areautomatically centered. When the gear is retracted, theends of the steering disconnect separate and the steeringwheel end locks in place. This ensures that the two endsengage properly when the gear is extended for landing.

6-7 LANDING GEAR POSITION AND WARNING SYSTEM

The nose gear and left and right main gear are allcontrolled by a single landing gear control handle. Thecontrol handle is part of a landing gear control panellocated on the right side of the center instrument panel.The control panel also has position indicators for each ofthe gear, a test button for the control handle warninglights and horn, a silencing button for the horn, and amanual release for the control handle lock. Severallanding gear position and lock-actuated switches areused in the landing gear control and warning circuits.Each main gear has a downlock switch, a pressuresequence switch, a bogie position switch, a pair oftouchdown switches, an uplock switch, and a flapperover-center lock switch.

Nose GearThe nose gear has an up switch, a lock switch, a positivedown switch, and an avionics switch. When the aircrafttakes off and the control handle is placed in the UPposition (Fig 6-9), the landing gear up relay is energized.The main gear downlock actuator releases the downlockand the main gear actuator retracts the gear. Hydraulic

pressure also causes the nose gear up/down lock actuatorto release the up/down lock and the nose gear actuator toretract the gear .

IndicationEach position indicator on the landing gear control paneldisplays the word UP when its corresponding gear is upand locked (Fig 6-9). A picture of a wheel appears whenthe gear is down and locked and diagonal stripes indicatethe gear is in transit. The main gear indicators arecontrolled through the downlock and uplock switches.The nose gear indicator is controlled by the up switchand lock switch for up position, and the positive downswitch and lock switch for the down position. A smallpanel to the left of the landing gear control panelcontains two bogie position indicators, one for eachmain landing gear. As the main gear extends, the axlebeam (bogie) is positioned perpendicular to the shockstrut. This actuates the bogie position switch and causesthe bogie position indicator to display a wheel. If thebogie is not in the proper position for landing, theposition indicator will show diagonal stripes.


6-13

Visual/Audio WarningRed warning lights in the landing gear control handlecome on whenever the handle and the gear are not in thesame position. The warning lights will also come on ifany of the throttles are retarded below minimum cruisepower without lowering the landing gear. This also setsoff a warning horn in the flight station. The horn can besilenced by pressing the HORN SILENCE button on the

landing gear control panel, but will sound again ifanother throttle is retarded past minimum cruise (Fig6-9). When the aircraft lands, its weight compresses themain gear shock struts and actuates a pair of touchdownswitches on each main landing gear. Actuating thetouchdown switches causes a solenoid in the landinggear control panel to lock the control handle in the DNposition. The control handle lock prevents the landinggear from being accidentally retracted while the aircraftis on the ground.


6-14

Fig 6-9, Landing gear control panel

1. BOGIE POSITION INDICATION PANEL2. LEFT MLG BOGIE POSITION INDICATOR3. RIGHT MLG BOGIE POSITION INDICATOR4. LANDING GEAR CONTROL PANEL5. LEFT MLG POSITION INDICATOR6. NLG POSITION INDICATOR7. RIGHT MLG POSITION INDICATOR8. UP AND LOCKED INDICATION9. INTRANSIT INDICATION

10. DOWN AND LOCKED INDICATION11. CONTROL HANDLE DOWN LOCK

RELEASE BUTTON

12. WARNING LIGHT AND HORN TEST SWITCH13. HORN SILENCE SWITCH14. CONTROL HANDLE WARNING LIGHTS15. BOGIE IN POSITION INDICATION16. BOGIE NOT IN POSITION INDICATION

SECTION VII FLIGHT CONTROLSYSTEM

This section consist of information extracted from T.O. 1C-141B-27GS-00-1.


1. The purpose of the flight control system

2. The function of

• primary flight control system

• aileron system

• rudder system

• elevator system

• secondary flight controls

• pitch trim system

• flaps system

• spoilers system

• flight control hydraulic system

• depleted uranium counterweights


• primary flight control system

• aileron system

• rudder system

• elevator system

• secondary flight controls

• pitch trim system

• flaps system

• spoilers system

• depleted uranium counterweights

Flight Control System

7-1


7-2

7-1 FLIGHT CONTROL SYSTEM

The flight control system of the C-141 aircraft includesthe primary and secondary flight controls, associatedcockpit controls, and the stall prevention system. Theoverall function and operation of the flight control

system is described in this section. A more detaileddescription of each subsystem is presented in subsequentsections of this manual.

7-2 PRIMARY FLIGHT CONTROL SYSTEM

The primary flight controls include the ailerons, rudder,elevators, and the associated power and controlcomponents of each (Fig 7-1). They are utilized formaintaining attitude and directional control of theaircraft. The ailerons are controlled by turning thecontrol wheel. The elevators are controlled by fore and

aft movement of the control column. The rudder iscontrolled by pushing the rudder pedals. Two completesets of controls are provided, one for the pilot and one forthe copilot. Either set can be used for aircraft control.Cockpit control output is transmitted to the powercontrol units through mechanical rods and cables.


7-3

Fig 7-1, Primary flight controls

Control CablesThree types of control cables are used. The first type isflexible twisted-steel-wire cable. A second type isLockclad cable. Lockclad is standard twisted-steel-wirecable with aluminum tubing swaged around it. This typeof cable has two advantages: the cable expands about thesame as the aircraft fuselage with temperature changes,and Lockclad cable is more rigid so there is less cable sagand fewer supports are needed for a given length ofcable. The third type of cable is nylon-jacketed cableused in the wing runs for the ailerons. Cable wear isreduced for nylon-jacketed cables. Vibrations are alsodampened out.

Tension RegulatorsEach primary flight control system has tensionregulators to compensate for changing temperatures (Fig7-2). The regulators are located under the flight deck.Each regulator consists of two metal quadrants which aremounted on a common shaft. The quadrants are

connected by a compression spring assembly. Theregulator takes in or lets out cable as required to maintainthe desired tension in the cables. A scale on eachregulator is used to rig the cable tension in relation toambient temperature.

Artificial FeelNormal power-on operation of the flight controls doesnot provide a resisting force, or feel, to indicate theamount of force required to displace the control surfacesinto the airstream. Without this feel force, there is atendency to over control the aircraft. The hinges andother structure are subjected to greater forces than arenecessary. Therefore, a double acting spring cartridge isattached between aircraft structure and each inputquadrant assembly to resist quadrant rotation. Thisresisting force is translated into an artificial feel felt bythe pilots when any control wheel, column, or rudderpedal is moved from its neutral position. The feel springcartridge also aids in returning the control surface(s) toneutral following movement of the surface.

7-3 AILERON SYSTEM

Roll motion of the aircraft is controlled by two ailerons.The ailerons attach to the outer wing rear beam. Theyform part of the trailing edges of the wings. The aileronsare simultaneously deflected up or down, but move inopposite directions to produce roll motion. Each aileronis normally actuated by a power control unit powered bythe No. 1 and No. 2 hydraulic systems (Fig 7-3). Thepower control unit is controlled by the pilot’s andcopilot’s control wheels. Movement of the controlwheels is transmitted through dual cable systems to acommon input quadrant assembly mounted on the rearbeam of the center wing. Pushrod and wing cablesystems then transmit motion from the input quadrant tothe power control units, which in turn actuate the aileroncontrol surfaces. An interconnect rod between the pilot’sand copilot’s tension regulators, and the common inputquadrants, serves to link up what are essentiallyindependent control systems. Automatic roll axis controlis provided by an autopilot servo connected to the inputquadrant by a closed loop cable.

ServotabsA servotab is hinged to the center section of each aileronrear beam (Fig 7-4). It forms the trailing edge of theaileron in this area. The tabs remain locked and fairedwith the ailerons during normal power-on operation. Ifnormal hydraulic power is lost, the tabs are unlocked andmove up or down in response to control wheel rotation .

Aileron Trim SystemRoll axis trim is accomplished by deflecting the aileroncontrol surfaces by means of an electrically-poweredactuator. The unit is connected to the input quadrantassembly at the center wing. Operation of self-centeringswitches on the center console actuates the trim unit. Atrim position indicator is located on the centerinstrument panel. Signals to the position indicator comefrom a position transmitter, which is part of the trimactuator.


7-4


7-5

Fig 7-2, Aileron controls


7-6

Fig 7-3, Aileron hydraulic components


7-7

Fig 7-4, Aileron system components

7-4 RUDDER SYSTEM

Directional control of the aircraft is provided by therudder. The rudder is hinged to the rear beam of thevertical stabilizer. The rudder is normally actuated by apower control unit powered by the No. 1 and No. 2hydraulic systems . The power control unit is controlledby the pilot’s and copilot’s rudder pedals (Fig 7-5).Movement of the rudder pedals is transmitted throughdual cable systems to a common input quadrantassembly located near the rudder yoke. Dual pushrodstransmit motion from the input quadrant to the powercontrol unit. In turn, the rudder control surface isdeflected left or right. An interconnect rod between thetension regulators and the common input quadrantsserves to link up what are essentially independentcontrol systems. Automatic directional control isprovided by two autopilot yaw damper servos. Theseunits are linked to the servo valve on the power controlunit.

Power Control Unit Hydraulic PressureOperating pressure of the rudder power control unitchanges as a function of airspeed. Reduced hydraulicpressure is used for high airspeeds to prevent excessivehinge loads during cruise. For low speed flight (below160 knots), faster response of the rudder is required.Thus, a higher operating pressure is used for low speedflight (Fig 7-6).

Rudder Trim SystemYaw axis trim is accomplished by deflecting the ruddercontrol surface by means of an electrically-poweredactuator. This unit is connected to the input quadrantassembly. Operation of self-centering switches on thecenter console actuates the trim unit. A trim positionindicator is located on the center instrument panel.Signals to the position indicator come from a positiontransmitter, which is part of the trim actuator (Fig 7-7).

7-5 ELEVATOR SYSTEM

Pitch attitude of the aircraft is controlled by twoelevators. The elevators are attached to the rear beam ofthe horizontal stabilizer to form the trailing edges of thehorizontal stabilizer. There is one elevator on each sideof the horizontal stabilizer bullet. The elevators aresimultaneously deflected up or down to produce anose-up or nose-down attitude of the aircraft. Theelevators are actuated by a power control unit mounted inthe stabilizer bullet. The power control unit is normallypowered by the No. 1 and No. 2 hydraulic systems. Theunit is controlled by the pilot’s and copilot’s controlcolumns. Movement of the control columns istransmitted through dual cable systems to a common

input quadrant located in the vertical stabilizer near thehorizontal stabilizer pivot. Dual pushrods transmitmotion from the input quadrant through an idlerbellcrank to the power control unit. The elevator controlsurfaces are in turn deflected by movement of the powercontrol unit actuating cylinders, output pushrods, andtorque tubes. An interconnect rod between the pilot’sand copilot’s tension regulators and the common inputquadrants serves to link up what are essentiallyindependent control systems. Automatic pitch axiscontrol is provided by an autopilot servo. It is connectedto the input quadrant by a closed loop cable.


7-8


7-9

Fig 7-5, Rudder system controls


7-10

Fig 7-6, Rudder system hydraulic components

Artificial Feel SystemArtificial feel forces an allowable elevator controlsurface travel change as a function of airspeed. Aservomechanism shifts the attach point between theelevator artificial feel cylinder assembly (on aircraft notmodified by TCTO 1C-141-716) or artificial feel andcentering assembly (on aircraft modified by TCTO1C-141-716) and the input quadrant. The shift toward oraway from the center of the quadrant changes the torquerequired to rotate the quadrant. The servomechanismitself is a linear actuator that responds to airspeed signalsfrom the central air data computers (CADCs). A manualoverride is provided to fix the feel force should a CADCmalfunction occur.

Loss of Normal Hydraulic PressureShould either the No. I or No. 2 hydraulic system

become inoperable, the elevators can be controlled withthe remaining system. However, the No. 3 hydraulicsystem and the power control unit emergency actuatingcylinder can be used with the remaining system torestore dual hydraulic system control. If both the No. 1and No. 2 hydraulic systems become inoperable, theelevators can still be controlled with only the No. 3hydraulic system.

Stall Prevention SystemThe stall prevention system monitors speed, angle ofattack, and other flight information. The systemcomputer determines the permissible angle of attack forthe existing flight conditions. When the permissibleangle of attack is exceeded, the system warns the pilotsby shaking the control columns and sounding an audiblewarning.


7-11

Fig 7-7, Rudder trim components

1. CONTROL SWITCHES2. POSITION INDICATORS3. FEEL SPRING CARTRIDGE

LEVEL ARM4. ARTIFICIAL FEEL SPRING

CARTRIDGE5. TRIM ACTUATOR (AND POSITION

TRANSMITTER)

7-6 SECONDARY FLIGHT CONTROLS

The secondary flight controls include the horizontalstabilizer (pitch trim), the wing flaps, the spoilers, andthe associated power and control components of each(Fig 7-8). Pitch trim is controlled electrically,hydraulically, or electro-hydraulically by variousswitches and/or control levers. The flaps are controlledby movement of the flap control handle. The spoilers arecontrolled by movement of the spoiler control handle.Depending upon the system, output from the cockpitcontrols is transmitted to the power control units throughwires, mechanical linkages, and cables.

Control CablesTwo types of control cables are used. The first type isflexible twisted-steel-wire cable. The second type ofcable is nylon-jacketed cable. It is used in the wing runs

for the spoilers. Cable wear is reduced whennylon-jacketed cables are used. Vibrations are alsodampened.

Tension RegulatorThe pitch trim subsystem is the only secondary flightcontrol subsystem with cables long enough to require atension regulator. The tension regulator compensates forchanging temperatures that lead to fuselage stretch orshrink. The regulator is located behind an access panelon the forward bulkhead of the cargo compartment. Theregulator takes in or lets out cable as required to maintainthe desired tension in the control cables. A scale on theregulator is used to rig the cable tension in relation to theambient temperature.


7-12

Fig 7-8, Secondary flight controls

7-7 PITCH TRIM SYSTEM

A pitch trim actuator attaches vertically between theforward portion of the horizontal stabilizer and thevertical stabilizer structure. The actuator is equippedwith independent rotating nut and rotating screw drives.Pitch trim is achieved by rotating the screw within thenut utilizing the screw drive (electric operation) orrotating the nut on the screw using the nut drive (eitherelectro-hydraulic or manual hydraulic operation)(Fig7-10). Either mode of operation will cause the actuator toextend or retract. A brake device holds one drive whilethe other is operating.

Manual Hydraulic andElectro-Hydraulic SystemsA hydraulic motor, a combination flow control andshutoff valve, and a control valve actuator are mountednear the base of the pitch trim actuator (Fig 7-10). Themotor drives the nut portion of the actuator through agearbox. The flow control and shutoff valve controlshydraulic fluid pressure to the motor and the direction offluid flow through the motor. In turn, the flow controland shutoff valve may be controlled either by the control

valve actuator and control wheel switches(electro-hydraulic system) or by cables and a lever(manual hydraulic system)(Fig 7-9).

Electric SystemAn electric motor drives the screw portion of the

actuator through clutches and a reduction gearbox. Theelectric motor is mounted near the top of the pitch trimactuator (Fig 7-10). Operation of the two switches on thecenter console engages the appropriate clutch forextension or retraction of the actuator. Autopilotelectrical signals also engage the appropriate clutch toprovide automatic control of pitch trim.

Travel and PositionAdjustable limit switches in the pitch trim subsystemlimit travel of the horizontal stabilizer in both directions.The limit switches and a position transmitter are locatednear the pivot hinges. Signals from the positiontransmitter are supplied to the position indicator locatedon the center instrument panel.

7-8 FLAP SYSTEM

Four double slotted Fowler-type flaps are located alongthe trailing edges of the wings. Two flaps are used perside (Fig 7-11). The flaps run from the wing root to theaileron. Extending the flaps changes the camber and areaof the wing into a high-lift configuration. Takeoff,approach, and landing speeds can then be reduced. Theflaps are mounted on carriages which roll on curvedtracks. The tracks extend aft from the trailing edge of thewing structure. Jackscrew actuators extend and retractthe flaps. Two hydraulic motors are mounted on agearbox attached to the center wing rear beam. Thegearbox is normally powered by the No. 2 and No. 3hydraulic systems and drives the jackscrews by means oftorque tubes. Flap movement is controlled by the flapcontrol handle located on the center console (Fig 7-11).Control cables transmit flap control handle movement toan input quadrant adjacent to the gearbox. The flaps andthe spoilers should not be deployed at the same time inflight. A high-force detent on the flap control handle

adds about 50 pounds to the pulling torque should thespoilers be opened (during flight). A position transmittermounted on the gearbox supplies signals to a positionindicator located on the center instrument panel.

Flap Asymmetry SystemAsymmetry brakes are installed at the ends of the torquetube run in each wing. The asymmetry brakes protectagainst the dangerous condition of uneven lift shouldopposite flaps extend or retract unequally. If either theinboard or outboard flap lags behind its symmetricalcounterpart on the opposite wing by more than threedegrees, the asymmetry system automatically actuatesboth asymmetry brakes on the torque tubes. This actioncloses the hydraulic shutoff valve on the flap drivegearbox. Since the asymmetry brakes can be reset onlyon the ground, the flaps are held in the position at whichthe asymmetrical condition began until the aircraft is onthe ground (Fig 7-13).


7-13

7-9 SPOILER SYSTEM

A total of 36 lift-spoiling panels are installed on theupper and lower surfaces of the aircraft wings (Fig 7-14).While in-flight, the spoilers are used as a speed brake toreduce speed for a high rate of descent. The spoilers mayalso be used to simply slow the aircraft down. On theground, the spoilers help shorten aircraft roll followinglanding by increasing the drag and decreasing the lift of

the wing. Each spoiler panel is hinged at its leading edgeto the wing. An actuator in each wing controls all spoilerpanels on that wing by a combination of pushrods,quadrants, and cables. The actuators are normallypowered by the No. 2 and No. 3 hydraulic systems.Spoiler movement is normally controlled by the spoilercontrol handle located on the center console (Fig 7-14), a


7-14

Fig 7-9, Pitch trim controls


7-15

Fig 7-10, Pitch trim drive unit

cable servo actuator below the control handle, andcontrol cables running to a center drive quadrant, then tothe actuators. The center drive quadrant is attached to thecenter wing rear beam. Two position transmitters (one ineach wing root) supply signals to a position indicatorlocated on the center instrument panel.

Spoiler Asymmetry SystemAn asymmetry detector is installed in each wing justinboard of the aileron control surface. If the outboardspoilers on one wing lag behind the outboard spoilers onthe opposite wing, the asymmetry system causes allspoiler panels to close. The spoiler system can bereturned to operational status following an asymmetrycondition by placing the spoiler control handle to theRESET position. This will not prevent the spoiler panelsfrom closing a second time should an asymmetrycondition reoccur.

Safety FeaturesA switch is provided on the center console adjacent tothe spoiler control handle for emergency closing of thespoilers. Automatic closing will also occur if the spoilercontrol handle is moved past the flight limit positionduring flight, or during a go-around condition. Reliefvalves incorporated in the spoiler servo actuators allowthe spoiler panels to move toward the closed position(blow down) in the face of excessive air loads. Thespoilers and flaps should not be deployed at the sametime in flight. A high-force detent on the spoiler controlhandle adds about 50 pounds of pulling force should theflaps be extended. The stall prevention and maximumspeed warning systems interface with the spoiler systemto provide a warning if spoiler operation is attempted atan airspeed and aircraft attitude bordering on stallconditions.


7-16


7-17

Fig 7-11, Flap control system components


7-18

Fig 7-12, Flap asymmetry system components


7-19

Fig 7-13, Flap asymmetry system wing components


7-20

Fig 7-14, Spoiler system components

7-10 FLIGHT CONTROL HYDRAULIC SYSTEMS

The flight controls are normally hydraulically actuated,except for pitch trim. Pitch trim may be eitherhydraulically or electrically actuated. When a powercontrol unit is powered by two hydraulic systems atonce, the control surface can be operated by either of thesupplying systems alone.

No. 1 Hydraulic SystemNo. 1 hydraulic system pressure goes to one actuator ineach aileron, elevator, and rudder power control unit.

No. 2 Hydraulic SystemNo. 2 hydraulic system pressure goes to one actuator ineach aileron, elevator, and rudder power control unit.

Pressure also goes to one wing flap motor, the hydraulicpitch trim motor, one wing spoiler power control unitactuator, and one-half of the tandem actuator spoilercable servo control.

No. 3 Hydraulic SystemNo. 3 hydraulic system pressure goes to one wing flapmotor, one-half of the tandem actuators in each wingspoiler power control unit, one-half of the tandemactuator spoiler cable servo control, one of the triplecylinder actuators in the elevator power control unit, andthe aileron tab lockout actuator.

7-11 DEPLETED URANIUM COUNTERWEIGHTS

WARNING - Failure to observe proper precautionswhen handling depleted uranium counterweightscould result in radiation exposure and illness.

Each aileron has a depleted uranium counterweight atthe outboard tip. Each elevator has depleted uraniumcounter weights to balance the control surface.

The counterweights are nickel coated and plated withcadmium. The cadmium plate is at least 0.005-inch thickto protect against corrosion. Radiation from the counterweights is negligible at a distance of two feet or more.Thus, the flight station and cargo compartment are notsubject to exposure from these counterweights. Observethe following precautions:

Do not machine, grind, file, weld, or burn thecounterweights. Small particles of uranium will igniteeasily.

Consult pertinent directives before starting rework in thearea next to the counterweights.

Personnel should touch the counterweights only toperform needed inspections and repairs.

Remove the counterweights in event of an accident.

Dispose of depleted uranium counterweights per TO00-110N-2.


7-21

C-141BPROGRAM 5362

AIRCRAFT SYSTEMS ANDPOWER PLANTS


Scott AFB, IL

PROGRAM 5362AIRCRAFT SYSTEMS AND POWERPLANT

TABLE OF CONTENTS

SECTION IPOWER PLANT

PARA TITLE PAGE

1-1 POWER PLANT . . . . . . . . . . . . . . . . . . . . 1-3

1-2 ENGINE . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-3

1-3 FUEL SYSTEM . . . . . . . . . . . . . . . . . . . . . 1-5

1-4 IGNITION SYSTEM . . . . . . . . . . . . . . . . . 1-8

1-5 AIR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . 1-9

1-6 CONTROL SYSTEM . . . . . . . . . . . . . . . . . 1-9

1-7 INSTRUMENT SYSTEM . . . . . . . . . . . . . 1-9

1-8 EXHAUST SYSTEM . . . . . . . . . . . . . . . . 1-13

1-9 OIL SYSTEM . . . . . . . . . . . . . . . . . . . . . . 1-16

1-10 STARTING SYSTEM . . . . . . . . . . . . . . . 1-19

SECTION IIFUEL SYSTEM

PARA TITLE PAGE

2-1 FUEL SYSTEM . . . . . . . . . . . . . . . . . . . . . 2-3

2-2 STORAGE SYSTEM . . . . . . . . . . . . . . . . . 2-3

Storage System Components. . . . . . . 2-3

Outboard Main Tanks . . . . . . . . . . 2-3

Outboard Main Tank Surge Box andEjectors . . . . . . . . . . . . . . . . . 2-3

Inboard Main Tanks . . . . . . . . . . . 2-3

Inboard Main Tanks Surge Box andEjectors . . . . . . . . . . . . . . . . . 2-4

Outboard Auxiliary Tank . . . . . . . . 2-4

Surge Box and Ejectors . . . . . . . . . 2-4

Inboard Auxiliary Tanks . . . . . . . . . 2-4

Inboard Auxiliary Tanks Surge Box andEjectors. . . . . . . . . . . . . . . . . . 2-4

Extended Range Tanks. . . . . . . . . . 2-4

Condensate Drain Valves . . . . . . . . 2-4

Fuel Vent System Components . . . . . 2-5

Fuel Vent Lines . . . . . . . . . . . . . 2-5

Fuel Vent Boxes . . . . . . . . . . . . . 2-5

Fuel Vent Box Seals . . . . . . . . . . . 2-5

2-3 DISTRIBUTION SYSTEM . . . . . . . . . . . . 2-9

Distribution System Components . . . . 2-9

Fuel Boost Pumps . . . . . . . . . . . . 2-9

Ejectors. . . . . . . . . . . . . . . . . . 2-9

Boost Pump Low Pressure WarningSwitches . . . . . . . . . . . . . . . . . 2-9

Fuel Level Control Valves . . . . . . . 2-10

Separation Valves. . . . . . . . . . . . 2-10

Crossfeed Valves . . . . . . . . . . . . 2-10

Engine Fuel Feed Lines . . . . . . . . 2-10

APU Fuel Supply Valve . . . . . . . . 2-10

iii

Emergency Fuel Shutoff Valves . . . . 2-10

Single Point Refueling (SPR) SystemComponents. . . . . . . . . . . . . . . 2-10

SPR Adapter and Adapter Housing. . . 2-11

Refueling Manifold . . . . . . . . . . . 2-11

Ground Isolation Valve . . . . . . . . . 2-11

Single Point Refueling Drain Pump andValve . . . . . . . . . . . . . . . . . . 2-11

Manual Drain Shutoff Valves . . . . . 2-11

Single Point Refueling ControlSwitches . . . . . . . . . . . . . . . . 2-11

2-4 AERIAL REFUELING SYSTEM (ARS)COMPONENTS . . . . . . . . . . . . . . . . . . . . 2-11

Universal Aerial Refueling ReceptacleSlipway Installation (UARRSI) . . . . 2-12

Aerial Refueling (A/R) Manifold andIsolation Valves. . . . . . . . . . . . . 2-12

Aerial Refueling Drain System . . . . . 2-12

Aerial Refueling Door Control Handle . 2-12

Aerial Refueling System Switches andLights . . . . . . . . . . . . . . . . . . 2-12

2-5 JETTISON SYSTEM . . . . . . . . . . . . . . . . 2-24

Jettison System Components . . . . . . 2-24

Jettison Valves . . . . . . . . . . . . . 2-25

Jettison Masts . . . . . . . . . . . . . . 2-25

Jettison Control Switches. . . . . . . . 2-25

2-6 INDICATION SYSTEM. . . . . . . . . . . . . . 2-25

Indication System Components. . . . . 2-26

Manifold Pressure Transmitter . . . . . 2-26

Manifold Pressure Indicator . . . . . . 2-26

Low Boost Pump PressureIndicators . . . . . . . . . . . . . . . . 2-26

Fuel In Temperature Indicator andSwitch. . . . . . . . . . . . . . . . . . 2-26

Fuel In Temperature Bulbs . . . . . . . 2-26

Fuel Quantity System Components. . . 2-26

Tank Units . . . . . . . . . . . . . . . 2-26

Compensators. . . . . . . . . . . . . . 2-27

Quantity Indicators . . . . . . . . . . . 2-27

Total Fuel Quantity Indicator. . . . . . 2-27

Thermistors . . . . . . . . . . . . . . . 2-28

Weep Hole Cracks . . . . . . . . . . . 2-29

SECTION IIIAUXILIARY POWER UNIT

PARA TITLE PAGE

3-1 AUXILIARY POWER UNIT (APU) . . . . . 3-3

3-2 APU ENGINE SYSTEM . . . . . . . . . . . . . . 3-3

3-3 APU FUEL CONTROL . . . . . . . . . . . . . . . 3-3

3-4 APU STARTING/IGNITIONSYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-3

Components . . . . . . . . . . . . . . . 3-3

Check Valves . . . . . . . . . . . . . . 3-3

Accumulators . . . . . . . . . . . . . . 3-5

Selector Valves. . . . . . . . . . . . . . 3-5

Selector Valve Switch . . . . . . . . . . 3-5

Flow Regulators . . . . . . . . . . . . . 3-5

Surge Accumulators . . . . . . . . . . . 3-7

Starter Motor . . . . . . . . . . . . . . . 3-7

Starter Adapter (Clutch) . . . . . . . . 3-12

Ignition Unit . . . . . . . . . . . . . . 3-12

Spark Ignitor Plug . . . . . . . . . . . 3-12

iv

3-5 APU BLEED AIR SYSTEM . . . . . . . . . . 3-12

Components. . . . . . . . . . . . . . . 3-13

Bleed Load and Flow Control Valve . . 3-13

Air Pressure Regulator . . . . . . . . . 3-13

Load Control Thermostat . . . . . . . . 3-13

3-6 APU CONTROL SYSTEM . . . . . . . . . . . 3-13

Components. . . . . . . . . . . . . . . 3-13

Bleed Load and Flow Control ValveSwitch. . . . . . . . . . . . . . . . . . 3-13

APU Door Control Toggle Switch . . . 3-13

APU Accumulator Selector ToggleSwitch. . . . . . . . . . . . . . . . . . 3-13

APU Control Switch . . . . . . . . . . 3-13

Exhaust Gas Temperature Gage . . . . 3-13

Annunciator Lights . . . . . . . . . . . 3-13

Emergency Control Panels . . . . . . . 3-13

3-7 APU OIL SYSTEM . . . . . . . . . . . . . . . . . 3-14

Components. . . . . . . . . . . . . . . 3-15

APU Oil Tank . . . . . . . . . . . . . 3-15

APU Oil Pump (Cluster) . . . . . . . . 3-15

Pressure Regulating Valve . . . . . . . 3-15

Oil Filter . . . . . . . . . . . . . . . . 3-15

Air-Oil Cooler . . . . . . . . . . . . . 3-15

Oil Pressure Switches. . . . . . . . . . 3-15

SECTION IVANTI-ICE AND RAIN REMOVAL SYSTEM

4-1 ICE AND RAIN PROTECTIONSYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3

4-2 WING ANTI-ICE AND EMPENNAGEDEICE SYSTEM . . . . . . . . . . . . . . . . . . . . 4-3

4-3 PITOT-STATIC TUBES ANTI-ICESYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3

4-4 WINDSHIELD ANTI-ICE AND RAINREMOVAL SYSTEM . . . . . . . . . . . . . . . . 4-3

4-5 ICE DETECTION SYSTEM . . . . . . . . . . 4-10

4-6 ENGINE PRESSURE RATIO PROBEANTI-ICE SYSTEM. . . . . . . . . . . . . . . . . 4-11

SECTION VHYDRAULIC SYSTEM

5-1 HYDRAULIC SYSTEM. . . . . . . . . . . . . . . 5-3

5-2 HYDRAULIC SYSTEM NO. 1 . . . . . . . . . 5-3

Supply System Components . . . . . . . 5-3

Reservoirs . . . . . . . . . . . . . . . . 5-3

Scupper Drain Valves . . . . . . . . . . 5-3

Suction Boost Pump . . . . . . . . . . . 5-3

Boost Pump Check Valve . . . . . . . . 5-3

Supply Shutoff Valves . . . . . . . . . . 5-3

Supply Line Quick DisconnectCoupling . . . . . . . . . . . . . . . . . 5-3

Pressure and Return SystemComponents . . . . . . . . . . . . . . . 5-3

Engine-Driven Pumps . . . . . . . . . . 5-3

High-Pressure Filters. . . . . . . . . . . 5-3

Pressure Line Quick DisconnectCoupling . . . . . . . . . . . . . . . . . 5-4

Pressure Shutoff Valves . . . . . . . . . 5-4

Isolation Check Valves . . . . . . . . . 5-4

System Relief Valve . . . . . . . . . . . 5-4

Pressure Indicating Components. . . . . 5-4

Return Check Valve . . . . . . . . . . . 5-4

Return Filter . . . . . . . . . . . . . . . 5-4

Drain System Components. . . . . . . . 5-4

v

System Operation . . . . . . . . . . . . 5-4

5-3 HYDRAULIC SYSTEM NO. 2 . . . . . . . . 5-10

Supply System Components . . . . . . 5-10

Reservoir . . . . . . . . . . . . . . . . 5-10

Scupper Drain Valves . . . . . . . . . 5-10

Suction Boost Pumps . . . . . . . . . . 5-10

Boost Pump Check Valve . . . . . . . 5-10

Supply Shutoff Valve. . . . . . . . . . 5-10

Supply Line Quick DisconnectCoupling . . . . . . . . . . . . . . . . 5-10

Pressure and Return SystemComponents. . . . . . . . . . . . . . . 5-10

Engine-Driven Pumps . . . . . . . . . 5-10

High-Pressure Filters . . . . . . . . . . 5-10

Pressure Line Quick DisconnectCoupling . . . . . . . . . . . . . . . . 5-10

Pressure Shutoff Valves . . . . . . . . 5-10

Isolation Check Valves . . . . . . . . . 5-10

System Relief Valve . . . . . . . . . . 5-11

Pressure Indicating Components . . . . 5-11

Return Check Valve . . . . . . . . . . 5-11

Return Filter . . . . . . . . . . . . . . 5-11

Drain System Components . . . . . . . 5-11

External Connections . . . . . . . . . . 5-11

System Operation . . . . . . . . . . . . 5-11

5-4 HYDRAULIC SYSTEM NO. 3 . . . . . . . . 5-12

Supply System Components . . . . . . 5-12

Pressure and Return SystemComponents. . . . . . . . . . . . . . . 5-14

Electrically-Driven Pumps . . . . . . . 5-14

Hand Pump . . . . . . . . . . . . . . . 5-14

Accumulators . . . . . . . . . . . . . . 5-14

Case Drain System Components . . . . 5-14

SECTION VIELECTRICAL SYSTEM

PARA TITLE PAGE

6-1 ELECTRICAL SYSTEM . . . . . . . . . . . . . . 6-3

6-2 GENERATOR DRIVE SYSTEM. . . . . . . . 6-3

6-3 AC GENERATION SYSTEM . . . . . . . . . . 6-6

6-4 DC GENERATION SYSTEM . . . . . . . . . 6-10

6-5 EXTERNAL POWER SYSTEM . . . . . . . 6-12

6-6 EMERGENCY GENERATIONSYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . 6-13

SECTION VIIOXYGEN SYSTEM

PARA TITLE PAGE

7-1 OXYGEN SYSTEM . . . . . . . . . . . . . . . . . . 7-3

7-2 CREW OXYGEN SYSTEM. . . . . . . . . . . . 7-5

7-3 TROOP OXYGEN SYSTEM . . . . . . . . . . . 7-6

7-4 PORTABLE OXYGEN BOTTLES . . . . . . 7-9

SECTION VIIIAIR CONDITIONING SYSTEM

PARA TITLE PAGE

8-1 AIR CONDITIONING SYSTEM. . . . . . . . 8-3

8-2 BLEED AIR SYSTEM . . . . . . . . . . . . . . . . 8-4

8-3 EQUIPMENT COOLING SYSTEM . . . . . 8-4

8-4 PRESSURIZATION SYSTEM. . . . . . . . . . 8-5

8-5 HEATING AND COOLING SYSTEM . . . 8-7

8-6 TEMPERATURE CONTROLSYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . 8-10

vi

SECTION I POWER PLANT

This section consist of information extracted from T.O. 1C-141B-71 GS-00-1


1. The purpose of the power plant

2. The function of

• engine

• fuel system

• ignition system

• air system

• control system

• instrument system

• exhaust system

• oil system

• starting system


• engine

• fuel system

• ignition system

• air system

• control system

• instrument system

• exhaust system

• oil system

• starting system

Power Plant

1-1


1-2

1-1 POWER PLANT

The C-141B is equipped with four axial flow TF33P7flat rated, forward fan engines. They have a 16-stagesplit compressor, an 8-can annular combustion chamber,and a 4-stage axial flow twin spool turbine. The enginesare enclosed in nacelles and are supported by pylons.

The pylons attach to the bottom surface of the wing.Each engine includes its own cowling, fuel system,ignition system, air system, controls system, instrumentsystem, exhaust system, oil system, and starting system.

1-2 ENGINE

The engine consists of the compressor, diffuser,combustion chamber, turbine, and accessory operatingsections (Fig 1-2). The compressor section consists ofthe low pressure (N1) and high pressure (N2)compressors. The N1 compressor is a two-stage largediameter fan, plus seven additional stages and is drivenby the second, third, and fourth stages of the turbine. TheN2 compressor has seven stages and is driven by the firststage of the turbine. The airflow through the enginedivides after the second compressor stage into primaryand secondary airflow. The primary airflow continuesthrough the engine. The secondary airflow is divertedinto the bifurcated duct which directs the flow aft to beexpelled through the fan nozzle. A fan duct seal systempressurizes the seals at the forward and aft end of each aft

cowl door preventing fan air leakage. The engine nacelleconsists of the engine and installed components, nosedome, nose cowl, forward cowl doors, bifurcated duct,aft cowl doors, and hot end assembly. Access doors areprovided in the forward cowl doors for engine and CSDoil tank servicing. The aft cowling contains access doorsand three blowout doors. Auxiliary air inlet doors areinstalled around the nose cowl to provide an additionalsource of air for ground operation and high powersettings at slow airspeeds. These doors are spring-loadedclosed and are opened by differential pressure betweenthe inlet duct and outside air. A thrust reverser withtarget type (clamshell) doors is installed in the hot endassembly.

Power Plant

1-3

Fig 1-1, Left and right side of engine

RIGHT SIDE OF ENGINE

LEFT SIDE OF ENGINE


1-4

Fig 1-2, Main accessory gearbox

1-3 FUEL SYSTEM

The engine fuel system receives pressurized fuel fromthe fuel tanks. The fuel pump consists of two stages. Thefuel enters the first stage of the fuel pump. The first stageuses an impeller type pump to increase the pressure forthe second stage. The second stage increases the pressureto a valve necessary to circulate fuel through the systemand into the combustion chamber. External plumbingroutes the fuel from the first stage to the second stage. Ade-icier heater, de-icier heater filter, and two differentialpressure switches are installed between the two pumpstages. The de-icier heater consists of a heat exchangerand an actuator (Fig 1-4). When selected by the flightcrew, the actuator is opened, allowing bleed air to heatthe fuel circulating through the heater. The de-icier filterremoves any contaminants present in the fuel. A fuelpump differential pressure switch is installed betweenthe inlet and outlet of the fuel pump first stage. Ifdifferential pressure drops below a set value, a PUMPOUT light will come on. Another pressure switch sensesdifferential pressure across the filter. If differentialpressure rises above a set value, a FIL BYPASS lightcomes on indicating the filter is bypassing fuel. Fuelre-enters the fuel pump at the second stage. The gear typepump increases pressure to approximately 1000 PSI.Pressurized fuel from the second stage of the fuel pumpis delivered to the fuel control. The fuel control metersfuel at pressures and flow rates required to obtain desiredengine operation. The fuel control automaticallyprovides a fuel starting schedule to allow the engine toaccelerate to a steady idle RPM without exceeding thetemperature limits or compressor surge limits. Duringflight, it maintains a constant turbine inlet temperaturefor each given throttle setting. It also prevents overtemperature and compressor surging duringacceleration, and prevents flameout during rapiddeceleration. The fuel control is driven by the N2compressor through the accessory gearbox. Anelectrically controlled fuel shutoff actuator is connectedto the fuel control. The actuator opens or closes a valve inthe fuel inlet passage (in the fuel control). A solenoidoperated fuel enrichment valve is controlled by the flightcrew to add extra fuel during cold weather starting (Fig1-3). The valve is energized depending on the type offuel used and the ambient temperature. A mechanicalpower lever connects the fuel control to the throttlelinkage. The throttles are the only input required by theflight crew to control fuel flow rates to the combustionsection. Metered fuel leaving the fuel control enters thefuel flow transmitter. The transmitter imparts a rotationon the fuel which is sensed by a turbine. The turbinedeflects against spring pressure sending an electricalsignal to be used for monitoring fuel flow. Fuel leavesthe transmitter and enters a fuel oil cooler. The fuel oilcooler is used to decrease the temperature of hot engine

oil with the cool fuel (Fig 1-5). Warmer fuel atomizeseasier. A pressurizing and dump (P & D) valve is used todetermine primary and secondary fuel f lowrequirements. It also drains the primary fuel manifoldduring engine shutdown. Fuel leaving the P & D valvetravels through a fuel manifold, with two concentrictubes, to the fuel nozzles. An inner tube carriessecondary fuel and the outer tube carries primary fuel.There are six fuel nozzles in each burner can, for a totalof 48 for each engine. The fuel nozzles spray primaryand secondary fuel into the combustion chambers. Thefuel is burned and expanded across the turbines.

Power Plant

1-5

Fig 1-3, Fuel system controls and indicators


1-6

Fig 1-4, Fuel system components on left side of engine

Power Plant

1-7

Fig 1-5, Fuel system components on right side of engine

1-4 IGNITION SYSTEM

The ignition system (Fig 1-6) ignites the fuel air mixturein the combustion chambers. A 4 joule and 20 joulesystem are contained in a single exciter. The 4 joulesystem is used for continuous ignition during adverseweather or flying conditions. The 20 joule system isintermittent and is used only during engine starts. The

exciter receives an AC and DC voltage from the aircraft.The voltages are filtered, rectified, and increased beforeexiting the exciter. High tension leads carry the voltageto spark ignitors located in burner cans No. 4 and No. 5.High voltage from the exciter ionizes a gap at the end ofthe ignitors creating a spark to ignite the fuel.


1-8

Fig 1-6, Ignition systems components

1-5 AIR SYSTEM

The air system extracts bleed air for anti-icing, cooling,and the aircraft pneumatic system. Bleed air is alsodumped overboard during certain engine operatingcycles. Bleed air is extracted from each side of thediffuser. Bleed air is classified as inside diameter air andoutside diameter air. Inside diameter air has a lowvelocity and high static pressure. Large particles areseparated leaving clean air. Outside diameter air has ahigh velocity and tends to retain heavier particles. Bleedair is also extracted by separate ducts for engineanti-icing (Fig 1-7) and nacelle anti-icing. Bleed air forengine anti-icing is routed forward to a flow regulatorand anti-icing valve. When the valve is opened air flowsinto the inlet guide vanes and nose dome. Warm airprevents ice from forming on these areas. Bleed air fornacelle anti-icing is also routed forward through a flowlimiter and regulator shutoff valve. Air is routed to achamber in the nose cowl and directed to the inlet lipthrough header tubes. Bleed air for both systems isexhausted into the engine inlet. Both systems areoperated at the same time from a single switch on thepilot’s overhead panel. In addition to anti-icing, bleed air

is used to aid in component cooling. Outside air entersthe nacelle through louvers in the bottom of the aft cowldoors. Air flows up and around engine components andis exhausted overboard through ejector ducts. Bleed airis injected, through nozzles, into the inlet of the ejectorduct. A low pressure area is created causing the outsideair to rush through the nacelle at a faster rate. A pressureswitch and ejector nozzle shutoff valve automaticallycontrol the bleed air to the nozzles. Aboveapproximately 19,000 feet, the pressure switch signalsthe shutoff valve to close. Below 19,000 feet, thepressure switch signals the shutoff valve to open.Compressor stalls are prevented by a compressorunloading system. The system reduces the pressure ratioacross the compressor by bleeding 12th stage airoverboard. Two bleed valves are mounted on thecompressor intermediate case. The left valve is a 4-inchbleed valve while the right is a 6-inch bleed valve. The4-inch bleed valve is normally spring-loaded closed andthe 6-inch bleed valve is normally spring-loaded open. Ableed control and bleed reset control are used to signalthe bleed valves when to open or close.

1-6 CONTROL SYSTEM

Two sets of four throttles are used to control enginepower requirements. Each throttle set can be movedindependently of the others. The throttles have tworanges, forward and reverse thrust. A series of pushrods,quadrants, pulleys, and cables transmit throttle levermovement to fuel control inputs (Fig 1-8). ATHROTTLE FRICTION knob is installed to regulatethe friction on both sets of throttles. A REVERSETHRUST LIMITER knob is used to limit the enginethrust in the reverse range. Mechanical linkage (Fig1-11) ties all throttles together so one adjustment sets allthrottles. A pylon mounted tension regulatorcompensates for changes in cable tension due to airframedeflections and temperature changes. The regulator also

contains a surge lock. The surge lock prevents regulatormovement if there is a sudden change in cable tensionlike a broken cable. The control system also includesprovisions for an engine emergency shutdown system(Fig 1-9). A FIRE PULL handle mechanically andelectrically shuts down the engine and its relatedsystems. The FIRE PULL handle is connected to amanual fuel shutoff valve on the wing front beam. Aseries of pulleys, cables, and pushrods connect the twocomponents. A cam on the FIRE PULL handle impelleralso actuates an emergency shutdown switch. Whenactuated, the switch energizes three relays to shut offfuel, hydraulic, electrical, and air systems.

1-7 INSTRUMENT SYSTEM

The engine instruments consist of five vertical scaleinstruments (VSI). They include indicators for enginepressure ratio (EPR), N1 RPM, N2 RPM, exhaust gastemperature (EGT), and fuel flow (Fig 1-10). An enginevibration indicating (EVI) system is also included. TheEPR indicator indicates the pressure ratio betweenengine inlet air pressure (PTO) and exhaust gas pressure(PT 7). A probe mounted on the inboard end of the pylondirects PTO air to a transmitter. Six probes, manifolded

together, in the engine exhaust directs PT 7 air to thetransmitter. The transmitter resolves the pressure ratiointo an electrical signal. The signal is fed to the EPRindicator. The N1 RPM indicator indicates the speed ofthe low pressure compressor fan. The indicator iscalibrated in percent RPM. A tachometer generatorsends signals to an engine data converter (EDC) beforebeing directed to the indicator. The N2 RPM indicatorindicates the speed of the high pressure compressor. The

Power Plant

1-9

N2 system works the same way as the N1 system. TheEGT indicator indicates the temperature of the gasesexiting the engine. Six dual junction thermocouples areinstalled in the engine exhaust section. As thethermocouples heat up, a small voltage is generated andsent to the EDC. The EDC conditions and amplifies thesignal for use by the indicator. The fuel flow indicatorindicates fuel flow to the combustion chamber in poundsper hour. The transmitter produces an electrical signalwhich is directed to the EDC. From the EDC, the propersignal is sent to the indicator. The EVI system consists of

two pickups and an amplifier indicator. The pickupssense low and high frequency vibrations and sendcorresponding signals to the amplifier indicator. Theindicator amplifies the signal and displays the result inmillimeters. A test button on each indicator testscontinuity of the indicator, wiring, and pickups. AFILTER switch and PICKUP switch provide control tothe type of inputs to the indicator.


1-10

Fig 1-7, Engine anti-icing components

Power Plant

1-11

Fig 1-8, Engine throttle linkage


1-12

Fig 1-9, Engine emergency shutdown system components

1-8 EXHAUST SYSTEM

The exhaust system, or hot end assembly, is used todirect exhaust gases aft providing thrust. A primaryexhaust nozzle directs engine exhaust gases aft. Fandischarge air is also directed aft through a duct aroundthe primary exhaust nozzle. Combined, the fan andengine exhausts produce the net engine thrust. Theexhaust system also deflects the exhaust forwardthrough a thrust reverser system. Each independentthrust reverser system operates through the throttles. Thesystem permits reverse thrust and is used during groundoperation only. Two target type(clamshell), thrustreverser doors are extended or stowed by hydraulicactuators acting through a mechanical linkage. Thethrust reversers are mechanically and electricallyactuated. Each thrust reverser system consists of a

hydraulic oil pump, filter, two actuators, a control valve,a flow regulator, a mechanical lockout, and indicatorlights. Oil for actuation of the system is taken from theconstant speed drive (CSD) reservoir. The fluid iscirculated through the thrust reverser system and thenported through the CSD oil cooler and back to thereservoir. The thrust reverser pump provides thepressure to open or close the thrust reverser doors. It alsocirculates oil through the system for cooling purposesand a positive lock when the system is stowed. Oil fromthe pump circulates through a filter to removecontaminants. The thrust reverser control valve (Fig1-11) directs fluid to the extend or retract side of theactuators. It is also used to help regulate CSD oil flow.One electrical input and two mechanical inputs to the

Power Plant

1-13

Fig 1-10, Engine instruments

1. EPR INDICATOR2. N1 RPM INDICATOR3. N2 RPM INDICATOR4. EGT INDICATOR5. FF INDICATOR

control valve are required to extend or retract the thrustreversers. The electrical input is provided by twosources. During initial deployment, a thrust reverserswitch in the throttle quadrant energizes the controlvalve. After the thrust reverser linkage unlocks, a notlocked relay de-energizes. The relay remainsde-energized until the thrust reverser stows and locks.Both mechanical inputs are provided through the throttlelinkage. A linkage rod attached to a bellcrank assemblyprovides one input. The other input comes from anactuator screw mounted on the 30” throttle control rod.The actuator screw moves a lockout lever permitting thecontrol valve to deploy the thrust reverser. A flowregulator is installed in the line between the control valveand the actuators. The regulator provides regulated flowto the actuators during thrust reverser deployment. It alsoprovides a snubbing action against air loads. Whenstowing the thrust reverser, the regulator provides freeflow back to the control valve. The actuators provide theactual force required to move the doors. They contain anorifice and check valve which permits cooling flow

when the thrust reverser is not being used. Target typethrust reverser doors are connected to the actuatorsthrough linkage. The linkage moves drive links aft andoutboard, opening the doors. The links move forwardand inboard during the door closing cycle. Links lockovercenter holding the thrust reverser doors closed. Ahook and striker arrangement is also used to aid theovercenter locking mechanism during highaerodynamic loads. A mechanical lockout connects tothe thrust reverser linkage (Fig 1-12). A hook assemblykeeps the throttles from being moved to FULL REVuntil the doors are fully open. Switches in the thrustreverser system provide input to the control andindication circuits. Two extended switches turn on theTHRUST REV EXTENDED light when the doors arefully open. Two retracted switches control a not lockedrelay. The relay controls the THRUST REV NOTLOCKED light and holding circuit. The light comes onany time the thrust reverser is unlocked. A THRUSTREVERSER PRESS light comes on when the thrustreverser control valve is pressurized.


1-14

Fig 1-11, Thrust pump, filter, and control valve

Power Plant

1-15

Fig 1-12, Thrust reverser mechanical lockout

1-9 OIL SYSTEM

The oil system provides lubrication to internal enginebearings. Oil is routed to the bearings under pressure. Ascavenge system picks up the lubricating oil and sends itback to an oil tank. A breather system vents compressorair leakage from bearing seals overboard. An oil tank isthe beginning and ending place for the oil. Oil leaves thesaddle type tank and enters the oil pump. The oil pumpcontains two gear type elements, the pressure element,and scavenge element. Both elements are driven by thesame shaft. Oil is pressurized by the pressure elementand sent to the oil filter. The filter removes anycontaminants. If a pressure differential builds across thefilter, a bypass valve opens, permitting oil to flow aroundthe filter. A pressure relief valve keeps oil pressurewithin limits. Oil passes through external tubes andinternal passages to small wire mesh filters before beingdirected to the bearings. Two indicators and two lightsare used to monitor the oil pressure system. An oilpressure transmitter, downstream of the filter, senses oilpressure and sends the results to an OIL PRESSindicator. Next to the transmitter is an oil temperaturebulb (Fig 1-15). The bulb’s resistance changes with oiltemperature and is measured by the OIL TEMPindicator. Two pressure switches control a LOW OILPRESSURE light. One switch senses oil pressure andprovides a ground when pressure drops below a setvalue. The other switch provides a ground when a setdifferential pressure develops across the oil filter. Eitherswitch will cause the light to come on when actuated. ALOW OIL QTY light monitors the usable oil remainingin the oil tank. A float type switch in the tank provides aground for the light. After oil lubricates the bearings,three scavenge pumps return the oil to the mainaccessory gearbox. Another pump returns oil from oneof the bearing compartments directly to the air-oilcooler. A fifth scavenge pump, which is the scavengeelement of the oil pump, scavenges oil from the gearboxand directs it to the air-oil cooler. The radiator typecooler transfers heat from the oil to the air (Fig 1-13). Athermostatic and pressure relief valve controls the flowof oil. During cool oil temperature, oil bypasses theair-oil cooler. As oil temperature increases, the valvecloses allowing oil to flow through the cooler. The lastitem the oil flows through before re-entering the oil tankis the fuel-oil cooler. The cooler is controlled similar tothe air-oil cooler. The heat from the oil is transmitted tofuel instead of air. A breather system vents sealpressurization air overboard. A rotary breather separatesheavy oil particles from the air before venting the airoverboard.


1-16

Fig 1-13, Oil cooling system components

Power Plant

1-17

Fig 1-14, Oil system pressure components


1-18

Fig 1-15, Oil system indicating components

1-10 STARTING SYSTEM

The starting system (Fig 1-16) uses pneumatic power torotate the engine for starting. Each engine is equippedwith a starting system consisting of a starter, startercontrol valve, starter duct, and the starting and ignitioncontrol system. The starter control valve regulates airpressure going to the starter. The control valve also actsas a shutoff valve when the starting system is not in use.The starter, when applied with pneumatic power, rotatesthe N2 compressor rotor through the gearbox until apredetermined starter or engine speed has been reached.At this time a cutout switch, located in the starter, is

actuated and closes the starter control valve. When thestarter control valve is closed, the starter is disengagedfrom the engine by means of a clutch. The startingsystem provides automatic operation of the starter andcontrol valve once the STARTER button has beendepressed. The FUEL AND START IGNITION switchcircuit provides control of the starting fuel and ignitionand has no direct control over the engine starter or startercontrol valve. A STARTER VALVE OPEN light comeson anytime the starter control valve opens.

Power Plant

1-19

Fig 1-16, Starting system components


1-20

SECTION II FUEL SYSTEM

This section consist of information extracted from T.O. 1C-141B-2-28GS-00-1


1. The purpose of the fuel system

2. The function of

• storage system

• distribution system

• jettison system

• indication system

• aerial refueling system components (ARS)


• storage system

• distribution system

• jettison system

• indication system

• ARS components

Fuel System

2-1


2-2

2-1 FUEL SYSTEM

The fuel system of the C141B aircraft provides forstorage, distribution, jettison, and indicating. TheStarlifter’s fuel system is designed to provide long range

mission capability with maximum reliability and safety.A continuous flow of fuel is maintained to the enginesunder all operating conditions.

2-2 STORAGE SYSTEM

Aircraft fuel supply is stored in ten tanks which arelocated inside the wing. Each wing has five tanks: twomain tanks, two auxiliary tanks, and an extended rangetank. One main tank and one auxiliary tank supply fuelfor the tank’s respective engines. An extended rangetank supplies fuel to the two engines on each wing. Tankboundaries, which begin near the fuselage and extend tothe wing tip bulkhead, are formed by the integrallystiffened upper and lower wing panels, front and rearwing beams, and the reinforced end bulkheads. Theintegral construction provides the greatest possiblecapacity with minimal weight. Each tank is designed towithstand the maximum loads to which it may besubjected, either on the ground or in the air, at anyallowable altitude or condition. The basic fuel supply isstored in the four main tanks and the four auxiliary tanks.The usable fuel capacity of the mentioned eight tanks is15,336 U.S. gallons. Full extended range tanks increaseusable fuel to 23,592 U.S. gallons. All tanks can bepressure refueled selectively or in unison through thesingle point refueling and aerial refueling systems, orthey can be filled individually through conventionalfiller ports in the upper wing surface. On the lower wingsurface tank bottoms are condensate drain valves. Eachpoppet type valve can be depressed with a drain tubeassembly to catch tank condensation. Tank protectionfrom structural damage, due to excessive internal orexternal pressures, is incorporated in the fuel ventsystem. A separate discussion of the fuel vent system isadded to this section.

Storage System Components

The storage system components consist of the fuel tanksand their related surge boxes, ejectors, access panels,and condensate drain valves (Fig 2-4). Ten fuel tanks arecontained within the aircraft wings: four main tanks, fourauxiliary tanks, and two extended range tanks. Tankboundaries consist of the upper and lower wing surfaces,end bulkheads, and front and rear main wing beams. Fuelstorage is within the wing boundaries, classifying thewings as wet wings. Tank design and layout in each wingis identical.

Outboard Main Tanks

Each outboard main fuel tank is a single compartmentwhich primarily provides fuel to the related engine (Fig2-2). Located in the wing outboard ends, the No. 1 maintank is located in the left wing, and No. 4 main tank in theright wing. Since the tanks are identical, furtherdiscussion will refer to the No. 1 main tank only. The No.1 main tank extends from the outboard end of the No. 1auxiliary fuel tank to the wing tip bulkhead, and betweenthe front and rear wing beams. The tank top and bottomare formed by the upper and lower wing panels. Eachoutboard main tank has a capacity of 1,233 US gallons.

Outboard Main Tank Surge Box andEjectors

A surge box is located in the forward outboard corner ofeach outboard main tank (Fig 2-1). The surge box is along narrow compartment of approximately 250 USgallons capacity. Divided into two compartments, thesurge box outboard compartment is called the surge boxsump. The surge box sump has a 37 US gallon capacityand houses the primary and secondary boost pumps. Thesurge box maintains a constant fuel supply at the boostpump inlet, preventing boost pump starvation duringdescent or roll conditions when tank fuel supply is low.One-way flapper valves located near the surge boxbottom allow fuel flow into the surge box and surge boxsump. Openings near the surge box top, allow fuel andvapor to flow from the box to the tank preventingdifferential pressure buildup. Three scavenge ejectorskeep surge box sump fuel levels full when either or bothboost pumps operate. One ejector is located in the surgebox inboard end, one ejector in the tank inboard end, andone in the tank outboard end.

Inboard Main Tanks

The No. 2 and No. 3 main tanks are located in the inboardforward corner of each wing. Both tanks are identical,except for the No. 2 main tank which supplies auxiliarypower unit (APU) fuel. The upper and lower wingpanels, the ribbed end bulkheads on three sides, and thefront wing beam form the tank boundaries. Each inboard

Fuel System

2-3

main tank has approximately a 2,169 U.S. galloncapacity.

Inboard Main Tank Surge Box andEjectors

Each inboard main tank surge box is a singlecompartment located in the tank aft outboard corner.Each surge box has a 120 gallon capacity and houses theprimary and secondary boost pumps. One-way flappervalves, located around the surge box base, allow fuelflow into the surge box. Openings near the surge box topallow fuel and vapor flow from the box to the tankpreventing differential pressure buildup. One scavengeejector, located inboard of the surge box, provides surgebox fuel supply when either or both boost pumpsoperate.

Outboard Auxiliary Tanks

The No. 1 and No. 4 outboard auxiliary tanks are locatedadjacent to and inboard of their corresponding maintanks (Fig 2-1, 2-3). Constructed like the main tanks, theoutboard auxiliary tank boundaries are formed by the topand bottom wing panels, ribbed end bulkheads, and frontand rear wing beams. Each outboard auxiliary tank hasapproximately a 2,570 U.S. gallon capacity.

Surge Box and Ejectors

The aircraft auxiliary tanks have open top surge boxesaround the primary boost pumps (Fig 2-1). Located inthe tank aft outboard corner, the surge boxes are highenough to prevent primary boost pump cavitation at lowfuel levels. Each outboard auxiliary tank secondaryboost pump is located in the tank aft inboard corner. Twoejectors supply fuel to each surge box during primaryboost pump operation.

Inboard Auxiliary Tanks

The No. 2 and No. 3 auxiliary tanks are located in thewing inboard ends behind their respective main tanks(Fig 2-1). Each inboard tank is a single compartmentformed by the upper and lower wing panels, ribbed endbulkheads on three sides, and the aft wing beam.Approximately 1,696 US gallons of fuel constitute eachinboard auxiliary tank’s capacity.

Inboard Auxiliary Tank Surge Box andEjectors

Similar to the outboard auxiliary tanks, the inboardauxiliary tanks, on aircrafts AF61-2779 and up, haveopen top surge boxes around the primary boost pumps.Located in the tank aft outboard corner, the surge boxesare high enough to prevent primary boost pumpcavitation at low fuel levels. Each inboard auxiliary tanksecondary boost pump is located in the tank aft inboardcorner. Two ejectors supply fuel to each surge boxduring primary boost pump operation.

Extended Range Tanks

Located at each wing midspan, the extended range tanksconsist of dual compartments for long range flights. Dueto tank location, use or non-use will minimally affectaircraft center of gravity. Tank boundaries are formed bythe upper and lower wing panels, the outboard auxiliarytank bulkhead, the inboard main and inboard auxiliarytank bulkheads, and the front and aft wing beams. Aninternal bulkhead, or baffle, divides each extended rangetank into two compartments. Located at the bulkheadbottom, four flapper valves permit fuel flow from theinboard compartment to the outboard compartmentonly. Openings at the bulkhead top allow free air andvapor passage in both directions for proper venting.Although each extended range tank consists of twocompartments, the tank functions as one unit. Both, tankfuel level control valve and the valve’s over the wingfiller are located in the inboard compartment. Fuel flowsinto the inboard compartment, through the flappervalves, and into the outboard compartment. Eachextended range tank has primary and secondary boostpumps located in the outboard compartment aft side. Inaddition to supplying the wing fuel manifold, the boostpumps operate two ejectors which transfer fuel from theinboard compartment to the outboard compartment. Nosurge box exists in either extended range tank resultingin a large amount of unusable fuel left in the tank.Motivated by primary boost pump flow from the inboardauxiliary tank, the inboard auxiliary tank ejector systemscavenges fuel from the outboard compartment.Scavenged fuel is then transferred to the inboardauxiliary tank surge box.

Condensate Drain Valves

On aircrafts AF61-2775 through AF63-8086, 12condensate drain valves exist on the fuel tank bottoms:two in each outboard main tank and one each in theremaining tanks. On aircrafts AF63-8087 and up, 16condensate drain valves exist on the fuel tank bottoms:three in each outboard main tank, two in each inboard


2-4

main tank and one each in the remaining tanks (Fig 2-4).Mounted flush with the bottom wing panels, thecondensate drain valves are located at tank low points.Depressing the condensate drain valve poppets with acondensate drain tube assembly catches tank drainage.

Fuel Vent System Components

The fuel vent system protects the aircraft fuel tanks fromexcessive internal or external pressures which couldcause structural damage. Refueling, defueling, aerialrefueling, engine feed, jettison, or altitude andtemperature changes create pressure differentials. Thevent system compensates for pressure differentials,regardless if the aircraft is airborne or on the ground,through open ended plumbing and vent boxes vented tothe atmosphere. Identical vent systems exist in bothwings. The fuel vent system consists of vent lines, ventboxes, and vent box seals.

Fuel Vent Lines

All fuel tanks vent through a fuel vent line with anupturned bellmouth inlet near each fuel tank inboardside. The inlets are directly below the wing skin uppersurface. The vent line opposite ends discharge into vent

boxes or vent box interconnecting lines. The vent lineopen ends are positioned so that fuel expansion will notflood the vent system, or allow fuel above the fuel levelduring aircraft climb or flight level attitude. Theoutboard main fuel tanks have wrap around vent lines toprevent large fuel quantities from entering the ventsystem during certain flight attitudes.

Fuel Vent Boxes

The fuel vent boxes are separate, sealed compartmentsin the outboard main fuel tanks and extended range fueltanks aft inboard corners. Vent lines interconnect thevent boxes. The outboard main tank vent box vents to theatmosphere by a standpipe. The outboard main fuel ventbox vents the outboard auxiliary and outboard maintanks. The extended range fuel tank vent box vents theinboard main and inboard auxiliary fuel tanks. Theextended range fuel tank vents to the interconnectingvent box line.

Fuel Vent Box Seals

On aircraft AF61-2775 through AF61-2779 andAF63-8078 and up, baffle type seals are installed on thevent line bellmouth openings where the lines discharge

Fuel System

2-5

Fig 2-1, Internal tank and dry bay access provisions

into the outboard main fuel tank vent box. The seals prevent fuel surges from discharging directly from the


2-6

Fig 2-2, Outboard main tank provisions

Fuel System

2-7

Fig 2-3, Outboard auxiliary tank provisions


2-8

Fig 2-4, Storage system condensate components

LEFT WING SHOWNRIGHT WING SIMILAR

1. TYPICAL CONDENSATEDRAIN VALVE

vent line to the vent box standpipe opening and then overboard. Seal installation fits snugly against the ventbox access doors.

2-3 DISTRIBUTION SYSTEM

The aircraft distribution system provides the plumbing,pumps, and valves for supplying fuel to the engines.System design permits single point refueling anddefueling, aerial refueling, crossfeed from any tank toany engine, and ground transfer from any tank to anyother tank. A fuel management panel at the flightengineer’s station controls most system operations.Distribution system components comprise most of thesystem components except for the tanks andindication/quantity components. However, specificdiscussions about single point refueling and aerialrefueling have been included to simplify the distributionsystem discussion.

Distribution System Components

Distribution system components and operationencompass refueling, defueling, fuel transfer, enginefeed, crossfeed, and jettison operations. However, singlepoint refueling and aerial refueling will be discussedlater in this section and jettison in another section. Thedistribution components discussed here are fuel boostpumps, ejectors, boost pump low pressure warningswitches, fuel level control valves, separation valves,crossfeed valves, engine fuel feed lines, APU fuel supplyvalve, and emergency fuel shutoff valves.

Fuel Boost Pumps

The aircraft has 20 boost pumps, two in each tank (Fig2-5). The outboard pumps in the main and auxiliarytanks are the primary pumps and the inboard pumpsare the secondary pumps. The extended range tankpumps are called inboard and outboard pumps. Theextended range pumps are located in the tank outboardcompartments. The main tank pumps are the same typepump and are interchangeable with each other, but theyare not interchangeable with the auxiliary and extendedrange tank pumps. The main tank pumps supplypressurized fuel to the engines. The output of one pumpis sufficient to supply one engine demand under allconditions. The pumps can supply fuel directly fromtank to engine, tank to crossfeed to engine, or acombination of both. Boost pumps are also used forjettison, defueling, and fuel transfer. Each boost pump iscomposed of a pumping element and a discharge orscroll housing. The extended range and auxiliary tankboost pumps have vortex breakers between the scrollhousings and their mounts. Vortex breakers smootheninlet fuel flow to the boost pumps. The scroll housing

facilitates plumbing and electrical wiring; thus, lockingthe pumping element into the scroll housingautomatically makes the necessary connections. Themain tank boost pump scroll housings have inlet screensand bypass valves. The bypass valves, normally closeddue to pressure within the boost pump, open when theboost pump is not operating to allow the engines tosuction feed from the main tanks. The output of bothmain tank boost pumps can be supplied to a commonmanifold. A one-way check valve, installed in eachboost pump output line, provides independent boostpump operation. Each boost pump has an individualcontrol switch located on the flight engineer’s fuelmanagement panel. The main tank boost pumps’maximum ratings are 22 psi at zero flow and 6 psi at23,700 PPH flow. The auxiliary and extended range tankboost pumps are rated at 17,000 PPH flow at 40 psi withsurges to 50 psi at zero flow.

Ejectors

Ejectors, sometimes called eductors, are jet pumps (Fig2-5, 2 of 6). The ejector system provides fuel transfer anytime the boost pumps are operating. The ejectors aresized to transfer approximately three times as much fuelas the boost pumps. Fuel transfer via the ejectors isdependent upon a flow through the ejectors to create asuction which picks up fuel. The flow is supplied by fueltapped off the boost pumps. The ejectors have twofunctions: to maintain a fuel supply to the boost pumpsin the fuel feed system and , to scavenge the vent boxes inthe fuel vent system. Check valves in the main tank fuellines prevent ejectors from destroying the vacuumneeded during suction feed to the engines.

Boost Pump Low Pressure WarningSwitches

A boost pump low pressure warning switch is connectedto each boost pump discharge line (Fig 2-5, 2 of 6). Theswitches are installed outside of the fuel tanks on the aftwing beam and can be removed and replaced withoutentering the tanks. The switches illuminate the PRESSLOW light on the fuel management panel. The switchesfor the main tank boost pumps are set to illuminate thewarning light when boost pump pressure outputdecreases to 4 psi. When main tank boost pump pressureincreases to 5 psi, the applicable warning lights will goout. The switches for the auxiliary and extended range

Fuel System

2-9

tanks will illuminate their respective lights when boostpump pressure output decreases to 25 psi. The warninglights will go out when auxiliary and extended rangeboost pumps pressure increases to 27 psi.

Fuel Level Control Valves

Each tank contains one fuel level control valvepositioned near the tank top (Fig 2-5, 3 of 6). The valve isa piston type valve. The valves govern the maximumlevel to which the tanks may be filled during single pointrefueling or fuel transfer. Each valve automatically shutsoff fuel flow to its respective tank when the fuel reaches apreset level. Fuel may be stopped at levels lower than fullby applicable switches on the flight engineer’s fuelmanagement panel which electrically control eachvalve. Each valve consists of a valve body, a primarypiston, a bleed shutoff valve, a float and solenoid, asecondary piston, bleed shutoff valve, float andsolenoid, and a thermal relief valve. In each tank, exceptthe extended range tanks, an orifice plate is installedbetween the valve and the fuel supply line. The orificeplate hole size determines fuel flow rate through thevalve. Orifice plates allow all of the tanks to reach thesame fuel level at approximately the same time whenrefueling all tanks simultaneously.

Separation Valves

Intertank fuel flow is controlled by three separationvalves called the center, left, and right separation valves(Fig 2-5, 4 of 6). The center separation valve, located inthe common manifold in the center wing dry bay, is amotor driven, butterfly type valve. The valve separatesthe manifold into halves and prevents an unbalancedcondition during crossfeed operation. The left and rightseparation valves are located in each extended rangetank outboard compartments. The valve bodies aremounted on the aft wing beam inside of the tank and thegear trains and actuators are mounted on the aft wingbeam outside of the tanks. Closed during flight, the leftseparation valve directs fuel from both outboardauxiliary tank boost pumps and the left extended rangetank to the No. 1 engine.

Crossfeed Valves

Four crossfeed valves corresponding to engine feed linesto the common manifold, provide engine fuel supplyfrom sources other than the main tanks (Fig 2-5, 5 of 6).The No. 1 and No. 4 crossfeed valves are located near theinboard ends of the No. 1 and No. 4 auxiliary tanks andthe No. 2 and No. 3 crossfeed valves are located in theinboard compartments of each extended range tank. Likethe left and right separation valves, the crossfeed valvesdo not incorporate thermal relief features. The four

crossfeed valves are individually controlled by fourCROSSFEED switches on the flight engineer’s fuelmanagement panel. The CROSSFEED switches (rotarytype) indicate an open valve when the switch flow linealigns with the panel flow line.

Engine Fuel Feed Lines

A series of fuel feed lines supply engine fuel. Main tankboost pump supply lines directly supply fuel to the maintanks’ associated engines through the emergency fuelshutoff valve, fuel strainer, and fuel heater. The maintank supply lines are connected to the common manifoldthrough the separation valves and interconnect lineswhich tie into the supply lines between the main tankboost pumps and the emergency fuel shutoff valve. Theauxiliary and extended range tank boost pumps areconnected directly to the common manifold. Theauxiliary tank boost pumps output is connected to thecommon manifold by a single line, while each extendedrange tank boost pump output is coerced by individuallines separated by a crossfeed valve in the commonmanifold.

APU Fuel Supply Valve

The APU fuel supply valve, located in the center wingdry bay, is a motor operated, sliding gate type valve (Fig2-5, 6 of 6 ). Installed in the supply line running from theNo. 2 main tank to the APU, the APU fuel supply valvecontrols APU fuel supply. The APU fuel supply valve isnormally closed, opening only when the APU isoperating. A single thermal relief feature is incorporatedin the valve to relieve pressure differentials between theAPU and No. 2 main tank.

Emergency Fuel Shutoff Valves

The emergency fuel shutoff valves, sometimes calledthe manual firewall shutoff valves, are mounted on thewing front beam (Fig 1-9). One valve is provided foreach engine. The shutoff valves are cable operated by thefire emergency control handles located on the engine firecontrol panel above the pilot’s center instrument panel.Pulling a control handle closes the correspondingshutoff valve. Pushing the handle in to the normalposition opens the shutoff valve. Any excessive pressurebuild-up on the engine side of any closed shutoff valvewill be relieved by a thermal relief feature in each valve.

Single Point Refueling (SPR) SystemComponents

Refueling operations are normally accomplishedthrough the single point refueling provisions (Fig 2-6).


2-10

The single point refueling system permits rapid refuelingof any individual tank or all tanks simultaneously. Singlepoint refueling system components consist of twoadapters and adapter housings, refueling manifold,ground isolation valve, drain pump and valve, twomanual shutoff valves, vent check valves, and controlswitches.

SPR Adapter and Adapter Housing

Located in the right aft wheel well, two refuelingadapters are provided for refueling operations.Connected to the refueling manifold, each adaptercontains a spring-loaded valve that is normally in theclosed position. The valve opens after the service hoseconnection to the adapter is made. Fuel flows through theadapter and valve to the 4-inch refueling manifold viatwo 3-inch lines. The plumbing associated with thesingle point refueling adapters allows single pointservicing to any or all tanks.

Refueling Manifold

The refueling manifold extends from the right aft wheelwell fairing up through the fairing and to a tee connectionin the center wing dry bay. At the tee connection, therefueling manifold connects to the common manifoldwhich supplies fuel to the individual tank refuel lines.

Ground Isolation Valve

The ground isolation valve is located in the center wingdry bay in the refueling line between the refuelingadapters and the wing fuel manifold. The GRD ISOVALVE switch on the flight engineer’s fuelmanagement panel opens and closes the ground isolationvalve. Refueling and defueling operations require theground isolation valve to be open; however, the valve isnormally closed to isolate the refueling manifold fromthe common manifold. A manual lever and the valve isprovided to open or close the valve when no aircraftpower is available. If the valve is manually closed afterpower has been removed from the aircraft, the valve willautomatically open when power is applied. The groundisolation valve contains no thermal relief provisions.

Single Point Refueling Drain Pump andValve

The single point refueling drain pump and valve islocated in the right wheel well under the single pointrefueling adapters. The 28 VDC pump is controlled bythe LINE DRAIN switch on the flight engineer’s fuelmanagement panel. The drain valve is wired in parallelwith the drain pump and is also controlled by the LINEDRAIN switch. Connected to the low point of therefueling adapter manifold, the drain pump and valve,when energized, removes residual fuel from therefueling manifold after refueling operation completion.Residual fuel is then pumped into the No. 3 main tank.

Manual Drain Shutoff Valves

Two manual drain shutoff valves are provided to drainthe refueling manifold and adapters if the drain pumpcannot be used. One manual drain valve is in a 3/8-inchline which extends from the refueling manifoldupstream of the drain pump and overboard in the lowerpart of the right wheel well fairing. Access to the valve isthrough a small access door in the right wheel wellfairing. This manual drain valve drains the refuelingmanifold from the drain pump to the ground isolationvalve. A second manual drain valve drains the manifoldsbetween the refueling adapters and the drain pump.Accessible from inside the wheel well, the drain valve islocated in a 3/8-inch line which extends from eachrefueling adapter to a tee connector and overboardthrough the right wheel well fairing.

Single Point Refueling Control Switches

Two switches on the flight engineer’s fuel managementpanel control single point refueling and draining. TheLINE DRAIN switch actuates the single point refuelingdrain pump and opens the drain valve to drain residualfuel from the single point refueling manifold afterrefueling or defueling. The GRD ISO VALVE switchenergizes the ground isolation valve open to allow fuelflow between the single point refueling adapters and thewing common manifold.

2-4 AERIAL REFUELING SYSTEM (ARS) COMPONENTS

The aerial refueling system components may be dividedinto two groups: components directly related to thereceiving and movement of fuel, and the controls andindicators of system operation. Components associatedwith fuel reception and movement consist primarily of auniversal aerial refueling receptacle slipway installation

(UARRSI), along with a refueling manifold, drainsystem, and two refueling isolation valves. Aerialrefueling controls and indicators consist of a doorcontrol handle, fuel management panel indicator lightsand switches, aerial refueling boom disconnect

Fuel System

2-11

switches, and pilot’s and copilot’s indicator lights (Fig2-7).

Universal Aerial Refueling ReceptacleSlipway Installation (UARRSI)

The UARRSI is installed in an external fairing on theaircraft top centerline just aft of the pilot’s station (Fig2-7, 2 of 4). A self-contained unit, the UARRSI consistsof a housing, a combination door and slipway, arefueling receptacle, a door actuating cylinder, a boomlatch cylinder, a solenoid operated hydraulic controlvalve, a manually operated hydraulic control valve, aninduction coil, signal amplifier, boom contact switch,and boom latch switch. The UARRSI provides a meansof refueling any or all aircraft fuel tanks from a boomtype tanker aircraft. System design and capacity allow900 U.S. gallons of fuel per minute at 55 psi boom nozzlepressure to fill all fuel tanks simultaneously.

Aerial Refueling Manifold and IsolationValves

The 4-inch aerial refueling manifold is routed under thefairing, external of the fuselage pressure skin from theUARRSI receptacle, to a Y fitting inside the center wingdry bay. A 3-inch diameter line is routed from each Yfitting branch through an aerial refueling isolation valveand connects each branch into the fuel system cross shipmanifold, one on each side of the cross ship separationvalve.

Aerial Refueling Drain System

The aerial refueling manifold drain system consists of amotor operated drain valve, a check valve, an ejector,and interconnect tubing. Located in the center wing drybay, the drain valve is mounted on a 1-inch diameter lineoff the aerial refuel manifold. The LINE DRAIN switchon the flight engineer’s fuel management panel controlsthe sliding gate type drain valve. The check valve ismounted on the center wing dry bay side of the No. 3main tank bulkhead and the ejector is mounted inside thetank. Drain tubing connects the motor operated valve tothe check valve and the check valve to the induced flowpart of the ejector. An aerial refueling manifold manual

drain is located in the fairing aft of the UARRSI. A largeaccess panel must be removed for manual drain valveaccess.

Aerial Refueling Door Control Handle

Located in the overhead trim directly over the flightengineer’s seat, the aerial refueling door control handle,when pulled, actuates a hydraulic actuating cylinderwhich opens the UARRSI slipway door. The handle hasa detent lock in both open and closed positions to preventinadvertent operation. The UARRSI slipway doormechanism is spring-loaded and the spring will open thedoor when the handle is pulled if aircraft hydraulicpressure to the UARRSI is unavailable (Fig 2-7, 3 of 4).

Aerial Refueling System Switches andLights

Located on the AERIAL REFUEL section of the flightengineer’s fuel management panel the MASTER switchprovides power for the aerial refueling system operation(Fig 2-7, 4 of 4). The toggle switch has two positions:ON and OFF. Placing the MASTER switch to ON willpower the aerial refueling signal amplifier, the UARRSIslipway lights, fairing lights, and wing leading edgelights dimming transformers. In addition, the MASTERswitch de-energizes the aerial refueling/throttle switchcontrol relay and controls. Also located on the AERIALREFUEL section of the flight engineer’s fuelmanagement panel is the MODE SELECT switch. Thetoggle switch has two positions: NORM and ORIDE. Inthe NORM position, the MODE SELECT switchfunctions as part of the power circuit from the AERIALREFUELING CONTROL circuit breaker through theMASTER switch and RESET switch to the signalamplifier. Placing the MODE SELECT switch toORIDE bypasses power from the signal amplifier andcompletes a circuit to the override relay and the amberOVERRIDE lights. The ORIDE position permitscontinued aerial refueling system operation in casenormal operation through the signal amplifier is lost.Energizing the override relay utilizes the normal circuitsthrough the UARRSI door actuated switches and theUARRSI contact and latch switches.


2-12

Fuel System

2-13

Fig 2-5, (1 of 6), Distribution system components


2-14


Fuel System

2-15



2-16


Fuel System

2-17



2-18


14. APU FUEL SUPPLY VALVE

15. ENGINE FUEL FEED LINE

16. EMERGENCY FUEL SHUTOFF VALVE

Fuel System

2-19

Fig 2-6, Single point refueling components locations


2-20

Fig 2-7, (1 of 4), Aerial refueling components locations

1. AERIAL REFUELING MANIFOLD

2. AERIAL REFUELING BOOM

DISCONNECT SWITCHES

Fuel System

2-21

Fig 2-7, (2 of 4), Aerial refueling components locations

3. UARRSI RECEPTACLE

4. DOOR RETAINER ASSEMBLY

5. AERIAL REFUELING DOOR

CONTROL BELLCRANK


MANUAL DRAIN VALVE



2-22

Fig 2-7, (3 of 4), Aerial refueling system components

Fuel System

2-23

Fig 2-7, (4 of 4), Aerial refueling system components

2-5 JETTISON SYSTEM

The fuel jettison system permits fuel to be jettisonedfrom a jettison mast in each wing so that aircraft grossweight may be reduced. Normally, jettisoning fuel fromthe auxiliary and extended range tanks is required, suchas in situations where aircraft gross weight needs to bereduced to normal landing gross weight. However,system design allows fuel to be jettisoned from any or allfuel tanks in an emergency. Fuel can be jettisoned fromthe main tank after the fuel in the extended range andauxiliary tanks has been jettisoned. Jettisoning isaccomplished by the fuel boost pumps pumping fuel intothe wing fuel manifolds, through the separation andjettison valves, and out the jettison masts.

Jettison System Components

The jettison system consists of fuel boost pumps,separation valves, jettison valves, jettison masts, andjettison control switches.

Fuel Boost Pumps

See section 2-3 for discussion of fuel boost pumps.

Separation Valves

See section 2-3 for a discussion of separation valves.


2-24

Fig 2-8, (1 of 2 ), Jettison system components locations

Jettison Valves

Located inboard of the jettison masts near each end of thefuel manifold, the jettison valves are mounted on thewing rear beam (Fig 2-8, 1 of 2). The valve bodies aremounted inside the fuel tanks and the gear trains andactuators are mounted outside the tanks on the wing rearbeams. Each valve also has an external lever for manualoperation and valve position indication. If the valve ismanually closed after power has been removed from theaircraft, the valve will automatically reopen when poweris supplied. Neither jettison valve has thermal relieffeatures.

Jettison Masts

Each wing has a jettison mast which is a 2-inch diameterline which terminates in the structural island between theflap and the aileron.

The mast extends aft of the trailing edge approximately 4inches so that all discharged fuel will clear the aircraftstructure. Installed over the end of each mast is a flamearrestor constructed of 20 mesh brass wire screen. Sincethe center separation valve is normally closed, the fuel inthe left wing discharges from the left jettison mast andfuel in the right wing discharges from the right jettisonmast.

Jettison Control Switches

The jettison valves are controlled by individual, guardedtoggle switches located in the lower corners of the flightengineer’s fuel management panel (Fig 2-8, 2 of 2).Placing the jettison control switches to JETTISONopens the two jettison valves.

2-6 INDICATION SYSTEM

Fuel system indication and warning are displayed on theflight engineer’s fuel management panel. Except for theemergency fuel shutoff valves, all fuel system functionsare controlled from the fuel management panel. The fuel

indicators and warning lights reflect fuel pressure, boostpump low pressure, fuel temperature, fuel sump low,fuel jettison, and quantity subsystems. The fuel quantity

Fuel System

2-25

Fig 2-8, (2 of 2), Jettison system components and locations

subsystem components and operation are discussedseparately at the end of this section.

Indication System Components

The fuel indication system is displayed on the flightengineer’s fuel management panel. Warning indicationsand associated components for fuel subsystemsdisplayed on the fuel management panel are themanifold pressure transmitter and indicator, low boostpump pressure indicators, and the fuel in temperatureindicators, switch, and bulbs.

Manifold Pressure Transmitter

Connected to the center wing dry bay fuel manifold onthe left side of the center separation valve is the manifoldpressure transmitter. The synchro type transmitterconsists of a case housing, a rotor, and a stator. Fuelmanifold pressure controls the rotor position. Thetransmitter and its associated indicator are used to checkthe individual fuel boost pump pressure output, as wellas fuel manifold pressure whenever the manifold ispressurized.

Manifold Pressure Indicator

Located on the flight engineer’s fuel management panelis the integrally lighted FUEL PRESS indicator.Graduated in 5 psi increments, the indicator reads from 0to 100 psi. The indicator contains a stator and rotorassembly with the indicator pointer connected to therotor shaft. The transmitter stator is electricallyconnected to the indicator stator.

Low Boost Pump Pressure Indicators

The fuel boost pump low pressure switches on the aftwing beam outside of the fuel tanks illuminate PRESSLOW lights on the flight engineer’s fuel managementpanel. The warning lights are located adjacent to theBOOST PUMPS control switches. Only one light isprovided for both boost pumps in each main tank, whileindividual lights are used for the auxiliary and extendedrange tank boost pumps. The main tank PRESS LOWlights will come on when boost pump pressure drops to 4psi, and auxiliary and extended range tank PRESS LOWlights will come on when boost pump output pressuredecreases to 25 psi.

Fuel In Temperature Indicator andSwitch

Located on the lower center of the flight engineer’s fuelmanagement panel is an integrally lighted fuel

temperature gage and an INB OUTBD selector switch.The FUEL IN TEMP gage displays the temperature offuel supplied to the engines. When the selector switch isin the INBD position, the indicator registers thetemperature of the fuel entering the No. 2 engine fuelmanifold. When the selector switch is in the OUTBDposition, the temperature of the fuel entering the No. 1engine fuel manifold is registered on the indicator. Thetemperature of fuel entering engines No. 3 and No. 4 isnot directly monitored because the fuel temperature forengine No. 3 should be approximately the same as forengine No. 2. The same relationship exists betweenengines No. 1 and No. 4 fuel temperatures. Fueltemperature is measured in degrees centigrade on a 70 to+ 150 scale.

Fuel In Temperature Bulbs

Two resistance type temperature bulbs are located in thefuel inlet lines to engines No. 1 and No. 2 at the left wingfront beam (Fig 2-10). The right wing fuel inlet lines donot have temperature bulbs since the fuel temperatureshould be approximately the same in both wings.Depending on the INBD/OUTBD selector switchposition, one of the two temperature bulbs is circuited tothe FUEL IN TEMP indicator on the fuel managementpanel. Changes in fuel temperature cause the resistanceof a sensing element within the bulb to change. Theresistance change is reflected by the indicator whichregisters fuel temperature in degrees centigrade.

Fuel Quantity System Components

The fuel quantity system measures the fuel amount ineach tank and the total amount of fuel in all tanks.Consisting of tank units, quantity indicators, a total fuelquantity indicator, and thermistors, the fuel quantitysystem displays fuel quantity in pounds.

Tank Units

Capacitor tank units, or probes, are mounted on circularmounting plates which bolt to the wing top surface. Eachoutboard main tank has 12 tank units, the inboard maintanks have five tank units each, the outboard auxiliarytanks each have six tank units, the inboard auxiliarytanks each have four tank units, and the extended rangetanks have seven tank units. The tank units consist of twotubes which form inner and outer electrodes that act as acapacitor. The tubes are insulated from each other by anend cap and a housing assembly, securing the tubes in afixed position. Fuel enters the area between the twotubes, allowing the fuel to seek the same tank fuel level.Since the tank units mount upright in the tanks, each tankunit capacity rating will vary depending on tank fuellevel. Tank units with compensators are located in each


2-26

tank. The number and location of the tank unitscompensate for fuel level changes. If the fuel levelincreases at the tank unit at one tank end, the fuel levelwill decrease at the tank unit at the opposite end of thetank. Thus, if the fuel level shifts to one end of the tank,the total tank capacitance will remain the same, notchanging the quantity measurement.

Compensators

A fuel compensator is located in each fuel tank andmounts on one of the tank units located at the lower endof each tank. The tank unit with compensator is installedin the main fuel tank surge boxes. In the auxiliary andextended range tanks, the tank unit with compensator isinstalled next to the outboard boost pumps. Thecompensators sense fuel density changes largelyresulting from temperature changes. Since fuel densityaffects fuel quantity, the compensator capacitancechanges to offset the error in the quantity indicatingsystem.

Quantity Indicators

One fuel quantity indicator is provided for each tank.Located on the flight engineer’s fuel management panel,the indicators provide a fuel quantity indication inpounds for each tank (Fig 2-9). Each indicator contains aphase sensitive amplifier, a bridge circuit, a two phaseinduction motor, and a power supply. A press-to-testswitch is located next to each indicator providingoperation checkout of the indicator.

Total Fuel Quantity Indicator

Located on the top center of the flight engineer’s fuelmanagement panel is a total fuel quantity indicator. Thegage indicates the total amount of fuel in all of the fueltanks. The gage dial indicates total fuel quantity inpounds times 10,000. A digital readout in the dial centerindicates total fuel quantity in pounds. Circuited to theindividual fuel tank indicators, the total quantityindicator also responds when an individual fuel tankindicator press-to-test switch is actuated.

Fuel System

2-27

Fig 2-9, Indication and quantity system

Thermistors

Each main tank surge box has a thermistor sensingelement. The thermistors are attached to a tank unitassembly so that the thermistors will be uncovered by

fuel when fuel levels drop to approximately 50 percentin the surge box. The low level circuits then actuate theSUMP LOW warning circuit. A SUMP LOW light onthe flight engineer’s fuel management panel will go onto indicate sump fuel quantity is low.


2-28

Fig 2-10, Indication and quantity system

WEEP HOLE CRACKS

Due to the fuel leak problems with the C-141B, we areshowing the location of the weep holes. Cracks haveoccurred due to stress and age of the aircraft. Cracks inthe weep hole area cause the skin to crack thus causingfuel to leak from the wing. There have been different

repairs to these areas to include; reaming out the holesand installing a boron patch over the cracked area. Fornow, these repairs have allowed extended airframeservice life.

Fuel System

2-29

Fig 2-11, Weep hole locations


2-30

SECTION III AUXILIARY POWER

UNIT

This section consist of information extracted from T.O. 1C-141B-2-49GS-00-1


1. The purpose of the auxiliary power unit (APU)

2. The function of the APU

• engine system

• fuel control system

• start/ignition system

• bleed air system

• control system

• oil system

3. The location of components for the APU

• engine system

• fuel control system

• starting/ignition system

• bleed air system

• control system

• oil system

Auxiliary Power Unit

3-1


3-2

3-1 AUXILIARY POWER UNIT (APU)

The APU system of the C-141B aircraft is an onboardgas turbine supplying pneumatic and electrical power forengine starting, operation of the air conditioning andelectrical system on the ground.

The APU system includes the APU engine, engine fuelcontrol, starting/ignition, bleed air, engine controls,indicating, and oil systems.

3-2 APU ENGINE SYSTEM

The APU engine is a pneumatic and shaft power gasturbine engine (Fig 3-1 and 3-2). It provides pneumaticpower as clean, compressed air for operation of aircraftmain engine starters and the aircraft air conditioningsystem. The engine delivers pneumatic power and shaftpower either simultaneously or independently. The APUis installed in the forward compartment of the left mainlanding gear pod and weighs approximately 260 pounds.Two large panels under the APU compartment, and asmall door on the compartment’s outboard side, provideengine access. The APU compartment also containslouver doors for engine air intake and an engine exhaustoutlet. The APU uses aircraft system fuel and is started

with a hydraulic starter. Electrical ignition power issupplied from the aircraft’s battery or an external powersource. Once the engine is started, it will automaticallyoperate at a steady speed. Basic theory and principles ofgas turbine engine operation apply to the APU engine.The compressor supplies mass air flow, while the singlecombustion can release heat energy which expands andaccelerates the air mass. The turbine section extracts therequired heat energy to drive the compressor, APUaccessories, and AC generator. Safe operating gastemperatures in the turbine section are automaticallyprotected by both the fuel control system and bleed aircontrol.

3-3 APU FUEL CONTROL

The APU fuel control system consists of plumbing andautomatically operated regulating components tomaintain a near constant power turbine operating speedunder varying output bleed air loads (Fig 3-3). The fuelsystem’s main control unit is a speed sensing governorwhich regulates the fuel/air mixture. Fuel is gravity fedfrom the No. 2 main tank. The automatic and electricallycontrolled APU fuel system provides the proper fuel

flow for starting, acceleration, and ON-SPEEDoperation with or without APU bleed air. The constantspeed governor has no fuel flow control until the APU isat or near ON-SPEED RPM (42,100 +/- 300). Theacceleration control valve provides the proper fuelmetering for starting, and acceleration to ON-SPEEDRPM.

3-4 APU STARTING/IGNITION SYSTEM

The APU starting system consists of a hydraulic startermotor. The motor is powered by two accumulators. Theaccumulators are charged from the aircraft’s No. 3hydraulic system. Hydraulic pressure sources includeelectrically driven pumps and a hand pump. A selectorswitch on the APU control panel allows single orcombined accumulators selection. Once the startingcycle begins, a spark igniter plug and ignition unit areused to ignite the fuel and air mixture. Depending uponatmospheric temperature, the hydraulic accumulator(s)selected for start charges the hydraulic motor for startingthe APU. A switch on the flight engineer’s panel causeshydraulic power from the accumulators to the starter

motor. The APU should accelerate to a self-sustainingspeed in 20 seconds or less.

Components

The APU starting system consists of check valves, twoaccumulators, selector valves, selector valve switch,flow regulator, surge accumulator, starter motor, andstarter adapter. A spark ignitor plug and an ignition unitare the only ignition system components.

Check Valves

The check valves are located between the hydraulicpressure source and the accumulators (Fig 3-5, 1 of 2).


3-3


3-4

Fig 3-1, APU engine system components

Check valves keep hydraulic pressure from bleeding offback through the pumps.

Accumulators

The APU starting system uses two 400 cubic-inch,piston-type accumulators (Fig 3-5, 1 of 2). Theaccumulators are located in the cargo compartment nextto the APU. Each accumulators capacity allows one startunder normal conditions.

Selector Valves

Selection of one or both accumulators is achieved by twoselector valves located downstream from eachaccumulator (Fig 3-5, 1 of 2).

Selector Valve Switch

A three-position rotary switch on the flight engineer’sAPU control panel controls selector valves selection(Fig 3-4).


3-5

Fig 3-2, APU engine system components


3-6

Fig 3-3, APU fuel system components

Flow Regulators

Down stream from the two selector valves, the flowregulator automatically limits hydraulic flow to 16gallons per minute (Fig 3-5, 1 of 2).

Surge Accumulators

The surge accumulator is a 10 cubic-inch accumulatordesigned to absorb the initial pressure surge applied tothe starter when the selector valve energizes open (Fig3-5, 1 of 2).


3-7

Fig 3-4, APU starting system controls

1. APU ACCUM SEL SWITCH


3-8

Fig 3-5, (1 of 2), APU starting system components


3-9

Fig 3-5, (2 of 2), APU starting system components


3-10

Fig 3-6, APU ignition system components


3-11

Fig 3-7, APU bleed air system components

Starter Motor

The hydraulic starter motor mounts on the APUaccessory drive (Fig 3-5, 2 of 2 ). The starter motor is aseries of pistons which convert hydraulic power tomechanical rotation of the output shaft.

Starter Adapter (Clutch)

The starter adapter mounts the APU accessory housingon the top inboard side (Fig 3-5, 2 of 2). The adaptercontains the clutch assembly and the starter drive pawls.

Ignition Unit

Mounted on the compressor plenum’s lower right side,the APU ignition unit is a capacitor-discharge type unit(Fig 3-6). During normal start, the system energizes onlyfor approximately 20 seconds.

Spark Ignitor Plug

Connected by a high-tension lead from the ignitor unit,the ignitor plug attaches to the combustor cap assembly(Fig 3-6).

3-5 APU BLEED AIR SYSTEM

The APU bleed air system is primarily controlled by thebleed load and control valve mounted on the turbine

plenum. The bleed shutoff and load control valvefunctions as a positive shutoff of bleed air, has a rate of


3-12

Fig 3-8, APU exhaust gas temperature components

1. THERMOCOUPLE

opening, and modulates to control bleed air. A butterflyvalve, a pneumatic valve actuator, rate control valve, andbleed control solenoid valve comprise the bleed load andflow control valve assembly. The bled air system alsoincludes an air pressure regulator and a pneumaticthermostat.

Components

The APU bleed air system consist of a bleed load andflow control valve, air pressure regulator, and apneumatic thermostat (Fig 3-7).

Bleed Load and Flow Control Valve

The APU bleed load and flow control valve assemblymounts on the turbine plenum and consist of a butterflyvalve, a pneumatic valve actuator, rate control valve,and bleed air solenoid valve.

Air Pressure Regulator

The air pressure regulator is a diaphragm type pressureregulator which regulates variable pressures from theAPU into a relatively fixed pressure.

Load Control Thermostat

Mounted on the turbine exhaust flange, a pneumaticthermostat connects to the control air pressure line.

3-6 APU CONTROL SYSTEM

The APU control system is electrical and automatic. Thecontrol panel, located at the flight engineer’s station,consists of three toggle switches, a rotary switch, onegage, and three annunciator lights (Fig 3-4 thru 3-9). TheAPU control system has two fire emergency controlpanels. One panel is mounted directly below the APUcontrol panel on the flight engineer’s panel. The secondemergency control panel is located in the cargocompartment aft of the forward crew door.

Components

The control system consist of a bleed load and controlvalve toggle switch, APU door control toggle switch,APU accumulator selector toggle switch, APU controlswitch , an exhaust gas temperature gage , threeannunciator lights, and two emergency control panels

Bleed Load and Flow Control ValveSwitch

The switch on the flight engineer’s panel energizes andde-energizes the bleed load and flow control valve (Fig3-10). The three speed switch arms the bleed air loadswitch.

APU Door Control Toggle Switch

Another toggle switch located on the flight engineer’spanel opens and closes the APU inlet door.

APU Accumulator Selector ToggleSwitch

The APU control panel’s third toggle switch is the APUaccumulator selector switch. The three-position switchmay be placed in any position, BOTH, No. 2 or No. 1, forstarting the APU (Fig 3-4 thru 3-9). Depending on thetemperature outside.

APU Control Switch

The APU control switch is a rotary type switch locatedon the flight engineer’s panel (Fig 3-4 thru 3-9). TheAPU control switch turns to three positions, OFF, RUN,and START, and is spring-loaded from the START toRUN positions.

Exhaust Gas Temperature Gage

Mounted on the flight engineer’s panel, the exhaust gastemperature gage closely resembles those used foraircraft engines.

Annunciator Lights

The APU control panel has three annunciator lights:NOT CLOSED, START, AND ON SPEED. The amberNOT CLOSED light goes on as soon as the APU inletdoors move out to the close position.


3-13

Emergency Control Panels

The APU control system has two fire emergency controlpanels. One panel is mounted directly below the APUcontrol panel on the flight engineer’s panel (Fig 3-9).The second is located in the cargo compartment aft of theforward crew door.


3-14

Fig 3-9, APU controls and indicators

3-7 APU OIL SYSTEM

The APU oil (lubricating) system is a self-contained,positive pressure, dry sump type oil system. The systemprovides lubrication, in the form of militaryspecification MIL-L-23699 oil, for all turbine andaccessory section gears and bearings. A tank, oil pump,pressure regulating valve, oil filter, air/oil cooler, andpressure switches comprise the APU oil system. The oilpump pumps oil from the tank through the oil filter tovarious lubricating points. A system relief valvemaintains desired pressure. A duplex scavenge pumpremoves oil from the accessory section and the turbineand then pumps the oil to the oil cooler and back to thetank.

Components

The APU oil system contains an oil tank, pump, pressureregulating valve, oil filter, air-oil cooler, and pressureswitches.

APU Oil Tank

A stainless steel tank is located on the lower left side ofthe APU’s compressor plenum (Fig 3-10). The tank ismounted at the lowest point in the lubrication system sothat when the APU is static, oil will gravity flow back tothe tank. Thus, oil will not flood the accessory case orturbine bearing area when the APU is static.

APU Oil Pump (Cluster)

The APU oil pump is located on the accessory section atthe APU’s lower left side (Fig 3-10). The pump sucks oilfrom the tank, then delivers it under pressure forlubricating the APU’s accessory gear and turbinesections.

Pressure Regulating Valve

Located downstream from the oil pump outlet is thepressure regulating valve. The valve relieves excessivepressure back to the pump inlet.

Oil Filter

A micronic paper filter is housed in the oil pumpdownstream from the pressure regulating valve (Fig3-10). All system pressure oil is normally filtered unlessthe filter clogs, and then the bypass valve will open tomaintain oil flow.

Air-Oil Cooler

Above the tank, the air/oil cooler is strap-mounted on thecompressor plenum (Fig 3-10). The air/oil cooler acylindrical type, air-oil heat exchanger. Airflow fromthe APU-driven fan located on the APU’s accessorysection is continuous while the APU is running.

Oil Pressure Switches

Three oil pressure switches are provided in the APUlubricating system: the centrifugal speed switches, lowoil pressure switch, and sequencing oil pressure switch(Fig 3-10 and 3-11).


3-15


3-16

Fig 3-10, APU oil system components


3-17

Fig 3-11, APU oil system components


3-18

SECTION IV ANTI-ICE AND RAIN

REMOVAL SYSTEM

This section consists of information extracted from T.O. 1C-141B-2-31GS-001.


1. The purpose of the ice and rain protection system

2. The function of

• wing anti-ice and empennage deice system

• pitot-static tubes anti-ice system

• windshield anti-ice and rain removal system

• ice detection system

• engine pressure ratio probe anti-ice system


• wing anti-ice and empennage deice system

• pitot-static tubes ant-ice system

• windshield anti-ice and rain removal system

• ice detection system

• engine pressure ratio probe anti-ice system

Anti-Ice and Rain Removal System

4-1


4-2

4-1 ICE AND RAIN PROTECTION SYSTEM

The ice and rain protection system of the C-141B aircraftincludes wing and empennage anti-ice, pitot staticanti-ice, windshield anti-ice, ice detection, and engine

pressure ratio probe anti-ice systems. The purpose of thesystems are to detect and control icing conditions on theaircraft.

4-2 WING ANTI-ICE AND EMPENNAGE DEICE SYSTEM

The wing anti-ice system uses bleed air from thecrosswing manifold. The bleed air routes through sixmodulating and shutoff valves to the diffuser ducts.Bleed air ejected from nozzles on the diffuser ductsmixes with air inside the wings and then routes throughpassages in the leading edge skin (Fig 4-1, 1 of 2). Themixed air is discharged through louvered vents in eachpylon and in each wing tip. The system is capable ofanti-icing three leading edge sections of each wing: onesection between the pylons and two sections outboard ofeach outboard pylon. The system does not provideanti-icing for each wing leading edge section betweenthe fuselage and the inboard pylon. Controls for thissystem is located on the pilot’s overhead panel (Fig 4-2).Ice protection is also not provided for the verticalstabilizer, but a de-icing system is incorporated in thehorizontal stabilizer leading edge. Empennage de-icingis provided by electrically heated metal elementsimbedded in the leading edge sections of the horizontalstabilizer. The horizontal stabilizer leading edge isdivided into eight sections, each containing two

shedding areas and three parting strips (Fig 4-3). Thede-icing and temperature controllers route AC power tothe shedding areas and parting strips. The electricalpower to the shedding areas is cycled on and off, andthe ice is blown off by the airstream. Electrical power tothe parting strips is applied continuously. Control of thesystem is maintained by a temperature controller, ade-icing controller, and temperature sensing elements inthe leading edge of the horizontal stabilizer. Thede-icing controller provides power to the sheddingareas. The temperature sensors in the shedding areasprovide overheat protection and signals the controller toshut off power whenever the heating elements exceed32°C. The temperature controller provides power to theparting strips. The total temperature sensor signals thecontroller to remove power when a temperature of -29°Cis reached. The parting strip temperature sensors signalthe controller to remove power when the parting stripheating elements exceed 32°C. In air temperaturesbelow -29°C, the controller removes power from bothparting and shedding areas.

4-3 PITOT-STATIC TUBES ANTI-ICE SYSTEM

Each pitot-static tube contains a head and a mast heatingelement for de-icing (Fig 4-4). The heating elements arecontrolled by the manually operated applicable PITOT

HEAT switch on the pilot’s overhead panel. When theheating elements are energized and heated, iceaccumulation is removed from the probe.

4-4 WINDSHIELD ANTI-ICE AND RAIN REMOVAL SYSTEM

Windshield anti-ice is provided by an electrical heatingsystem (Fig 4-5). Three separate circuits control thewindshield anti-icing. Electrically conductive film onthe windshields acts as a heater element to deice anddefog the windshield. A transformer located in the rightavionics bay provides voltage to the windshield film.Control switches on the pilot’s overhead panel providedefogging and de-icing operation. The side windshieldshave defrost and defogging capabilities only. The pilot’sand copilot’s windshields are equipped with a jet blastrain removal system (Fig 4-6). Continuous slot-typenozzles supply high temperature, high velocity air at thebases of the windshields. Bleed air is supplied to the

nozzles from a point downstream of the air conditioningsystem primary heat exchangers. The air temperature iscontrolled to a maximum 460°F by the primary heatexchangers. The rain removal system is manuallyoperated with a selection switch on the pilot’s overheadpanel. The selection switch has four positions: OFF,PILOT, BOTH, and COPILOT. When the system is off,the pressure regulator shutoff valves and the nozzleshutoff valves are closed. When the selection switch isplaced in any of the remaining three positions, bothpressure regulator valves are opened when air pressure isapplied. The electric motor driven nozzle shutoff valveswill open, depending on the position selected.


4-3


4-4

Fig 4-1, (1 of 2),Wing anti-ice system components


4-5

Fig 4-2, (2 of 2),Wing anti-ice system components


4-6

Fig 4-3, Empennage de-ice system


4-7

Fig 4-4, Pitot static tubes anti-ice system


4-8

Fig 4-5, Windshield anti-ice system


4-9

Fig 4-6, Windshield rain removal system

4-5 ICE DETECTION SYSTEM

Icing conditions in the ambient air are monitored by anice detector in the left forward fuselage (Fig 4-7). The icedetector has two probes, one senses temperature and onesenses humidity, which protrude into the air streamflowing around the airplane. When conditions which

would cause ice formation on the engines, windshields,and leading edges are detected, a warning light comeson, and the engine anti-ice and engine pressure ratioprobe anti-ice systems automatically activate. Theremaining anti-ice systems must be manually activated.


4-10

Fig 4-7, Ice detection system

4-6 ENGINE PRESSURE RATIO PROBE ANTI-ICE SYSTEM

Integrated in each of the four engine pressure ratio (EPR)probes are anti-ice heating elements (Fig 4-8). Threeswitches, ENGINE ANTI-ICE, PITOT HEAT PILOT,and ICE DET CONTROL, are capable of activating theEPR probe anti-ice system. The EPR probe anti-ice

system will automatically come on if the ICE DETCONTROL switch is in AUTO and icing conditions aredetected. When the probes heating elements areenergized and heated, ice accumulation is removed fromthe probes.


4-11

Fig 4-8, Engine pressure ratio probe anti-ice system


4-12

SECTION V HYDRAULIC SYSTEM



1. The purpose of the hydraulic system

2. The function of

• hydraulic system No. 1







Hydraulic System

5-1


5-2

5-1 HYDRAULIC SYSTEM

The primary purpose of the hydraulic system is to powerthe aircraft’s hydraulically actuated systems andcomponents. The C-141 aircraft uses three separate andindependent 3000 PSI hydraulic systems: hydraulic

system No. 1, hydraulic system No. 2, and hydraulicsystem No. 3. Most components are in service centers inthe cargo compartment.

5-2 HYDRAULIC SYSTEM NO. 1

Hydraulic system No. 1 is powered by two variablevolume hydraulic pumps. They are connected in paralleland driven by engines No. 3 and No. 4. The function ofsystem No. 1 is to supply power to: one of the two aileronpower control unit cylinders, one of the two rudderpower control unit cylinders, and one of the threeelevator power control unit cylinders. A hydraulicreservoir and most components are in a service center onthe right side of the cargo compartment. The hydraulicsystem is operated from a hydraulic control panel at theflight engineer’s station (Fig 5-1).

Supply System ComponentsThe supply system provides hydraulic fluid underpressure to the engine-driven pumps. The supply systemcontains a reservoir, scupper and drains, suction boostpump, boost pump pressure switch, boost pump checkvalve, two supply shutoff valves, and a supply line quickdisconnect coupling.

ReservoirsA non-pressurized reservoir is mounted on the wall at theservice center. When properly serviced, the reservoircontains approximately 2.4 gallons of hydraulic fluid(Fig 5-2).

Scupper Drain ValvesA dish-shaped scupper and two drain valves dumphydraulic fluid through an overboard drain line (Fig 5-2).

Suction Boost PumpA suction boost pump is mounted below and slightlyforward of the reservoir (Fig 5-2). This provides anadequate supply of hydraulic fluid to the No. 3 and No. 4engine-driven pumps.

Boost Pump Check ValveMounted next to and downstream from the boost pumppressure switch is a one-way check valve (Fig 5-2). Thecheck valve holds a prime or head of hydraulic fluid tothe engine-driven pumps.

Supply Shutoff ValvesAfter leaving the check valve, the supply line runs alongthe leading edge of the wing and splits, one supply linegoing to each engine. Mounted to the wing leading edgeabove each engine is a supply shutoff valve (Fig 5-3).

Supply Line Quick Disconnect CouplingA quick disconnect coupling is provided in theengine-driven pump supply line. It is at the forward edgeof the pylon (Fig 5-3).

Pressure and Return SystemComponentsThis system provides the force to move cylinders in theprimary flight control power control units. The systemcontains engine-driven pumps, high-pressure filters,pressure line quick disconnect coupling, pressureshutoff valves, pressure switches, isolation checkvalves, system relief valve, pressure indicatingcomponents, and a return filter.

Engine-Driven PumpsAn engine-driven pump is mounted on the accessorygearbox of engines No. 3 and No. 4 (Fig 5-4).

High-Pressure FiltersA high-pressure filter is installed downstream of theengine-driven pumps on the right side of each engine(Fig 5-4). The filter element is made from corrosionresistant steel, and filters contaminants from the pump.

Hydraulic System

5-3

Pressure Line Quick DisconnectCouplingA quick disconnect coupling is provided in the pressureline at the forward edge of the pylon. This coupling is thesame type of coupling used in the supply line (Fig 5-3).

Pressure Shutoff ValvesMounted near each supply shutoff valve is a pressureshutoff valve (Fig 5-3).

Isolation Check ValvesAn isolation check valve is installed in the pressure lineof each engine. These check valves serve two purposes:to prevent reverse flow through an inoperativeengine-driven pump and to prevent the pressure fromone engine-driven pump from actuating the pressureswitch of the other engine-driven pump in its system.

System Relief ValveA system relief valve, mounted in the service center,protects the system from excessive pressure if anengine-driven pump fails to compensate pressure (Fig5-2).

Pressure Indicating ComponentsPressure out of both engine driven pumps can be read atthe hydraulic control panel or the service center (Fig5-2).

Return Check ValveWhen the system is depressurized, the check valveremains seated and prevents fluid from the empennagefrom draining into the reservoir (Fig 5-2).

Return FilterA system return filter installed next to the reservoir,removes contaminants returning from the primary flightcontrol power control units (Fig 5-2).

Drain System ComponentsCase drain flow from the engine-driven pumps is routedback to the reservoir. The system also provides propercooling and lubrication for the pumps (Fig 5-4).

System OperationThe flow of hydraulic fluid begins and ends at thereservoir. Fluid gravity feeds to the suction boost pump.The boost pump supplies fluid under pressure to bothengine driven pumps.


5-4

Fig 5-1, Hydraulic system No. 1 controls

Hydraulic System

5-5

Fig 5-2, Hydraulic system No. 1 service center components


5-6

Fig 5-3, Right wing and pylon hyd sys No.1 components

Hydraulic System

5-7

Fig 5-4, No. 1 system engine mounted components


5-8

Fig 5-5, Hydraulic system No.2 service center components

Hydraulic System

5-9

Fig 5-6, Left wing and pylon hyd sys No. 2 components


Hydraulic system No. 2 is powered by two variablevolume hydraulic pumps connected in parallel anddriven by engines No. 1 and No. 2. System No. 2 supplieshydraulic power to each of the following systems:ailerons, rudder, elevators, pitch trim, emergencygenerator, landing gear including steering and brakes,spoilers, flaps, and the universal aerial refuelingreceptacle slipway installation (UARRSI). A hydraulicreservoir and most of the components are located in aservice center on the left side of the cargo compartment.Provisions are made for interconnecting hydraulicsystem No. 3 with hydraulic system No. 2.

Supply System ComponentsThe supply system provides hydraulic fluid underpressure to the engine-driven pumps. The supply systemcontains a reservoir , scupper and drains,electr ical ly-dr iven suct ion boost pump,hydraulically-driven suction boost pump, boost pumpcheck valve, boost pump pressure switch, two supplyshutoff valves, and a supply line quick disconnectcoupling.

ReservoirA non-pressurized reservoir is mounted on the wall at theservice center (Fig 5-5). When properly serviced, thereservoir contains approximately 4.2 gallons ofhydraulic fluid with the landing gear down andapproximately five gallons with the landing gear up.

Scupper and Drain ValvesA dish-shaped scupper and two drain valves dumphydraulic fluid through an overboard drain line (Fig 5-5).

Suction Boost PumpsTwo suction boost pumps are provided for hydraulicsystem No. 2. The boost pumps provide an adequatesupply of hydraulic fluid to the No. 1 and No. 2engine-driven pumps (Fig 5-5).

Boost Pump Check ValvesMounted near and downstream from each boost pump isa one-way check valve. The check valve holds a prime,or head of hydraulic fluid to the engine-driven pumps(Fig 5-5).

Supply Shutoff ValvesAfter leaving the pressure switch, the supply line runsalong the leading edge of the wing. Mounted to the wingleading edge above each engine is a supply shutoff valve(Fig 5-6).

Supply Line Quick Disconnect CouplingA quick disconnect coupling is provided in theengine-driven pump supply line. It is at the forward edgeof the pylon (Fig 5-6).

Pressure and Return SystemComponentsThis system provides the force to operate systemsconnected to the hydraulic system. The system containsengine-driven pumps, high-pressure filters, pressureline quick disconnect coupling, pressure shutoff valves,pressure switches, isolation check valves, system reliefvalve, pressure indicating components, return checkvalve, and a return filter.

Engine-Driven PumpsAn engine-driven pump is mounted on the accessorygearbox of engines No. 1 and No. 2 .

High-Pressure FilterA high-pressure filter is installed downstream of theengine-driven pumps on the right side of each engine.

Pressure Line Quick DisconnectCouplingA quick disconnect coupling is provided in the pressureline at the forward edge of the pylon. This coupling is thesame type of coupling used in the supply line (Fig 5-6).

Pressure Shutoff ValvesMounted near each supply shutoff valve is a pressureshutoff valve (Fig 5-6).

Isolation Check ValvesAn isolation check valve is installed in the pressure lineof each engine. These check valves serve two purposes:


5-10

to prevent reverse flow through an inoperativeengine-driven pump and to prevent pressure from oneengine-driven pump from actuating the pressure switchof the other engine-driven pump in its system (Fig 5-6).

System Relief ValveA system relief valve, mounted in the service center,protects the system from excessive pressure if anengine-driven pump fails to compensate pressure (Fig5-5).

Pressure Indicating ComponentsPressure output of both engine-driven pumps can be readat the hydraulic control panel or the service center (Fig5-5).

Return Check ValveWhen the system is depressurized, the check valveremains seated and prevents fluid in the empennagefrom draining into the reservoir.

Return FilterA system return filter, installed next to the overflow tankassembly removes contaminants returning from theusing sub-systems (Fig 5-5).

Drain System ComponentsCase drain flow from the engine-driven pumps is routedback to the reservoir. The system also provides propercooling and lubrication for the pumps. The case drainflow system consists of a case drain line quickdisconnect coupling, relief valves, bypass check valves,and a case drain filter (Fig 5-5).

External ConnectionsGround test connections in the forward end of the leftmain landing gear wheel well allow an externalhydraulic power source to pressurize hydraulic systemNo. 2. Two connections are used: one for the supply lineand one for pressure (Fig 5-7).

System OperationThe flow of hydraulic fluid begins and ends at thereservoir. Fluid gravity feeds to the electrically-drivensuction boost pump. The boost pump supplies fluidunder pressure to both engine-driven pumps.

Hydraulic System

5-11

Fig 5-7, Hydraulic system No.2 external groundconnection


Hydraul ic system No. 3 is powered by twoelectrically-driven, variable-volume pumps. They areconnected in parallel. Hydraulic system No. 3 is used tosupply power to each of the following systems: spoilers,flaps, cargo doors, and APU starter. Hydraulic systemNo. 3 is also used for emergency operation of the brakes,elevators, and aileron tabs. The hydraulic reservoir andmost of the components are in a service center. Theservice center is on the left side of the cargocompartment just forward of the No. 2 system servicecenter. Hydraulic system No. 3 may also be used topressurize the No. 2 hydraulic system through an

interconnect valve. Control of the system is operatedfrom the No 3 system control panel, on the flightengineer’s hydraulic control panel.

Supply System ComponentsThe supply system provides hydraulic fluid undergravity flow to the electrically driven pumps. The supplysystem consists of a reservoir and a manual drain valve.The nonpressurized reservoir is mounted on the wall atthe service center. The reservoir contains approximatelyfive gallons of hydraulic fluid.


5-12

Fig 5-8, No. 3 system electrically driven pumps

1. PUMP NO 2 (behind panel)2. PUMP NO 1 CASE DRAIN LINE3. PUMP NO 1 PRESSURE LINE4. PUMP NO 1 SUPPLY LINE5. PUMP NO 2 SUPPLY LINE6. PUMP NO 2 PRESSURE LINE7. PUMP NO 2 CASE DRAIN LINE8. PUMP NO 1 (behind panel)

Hydraulic System

5-13

Fig 5-9, No. 3 System service center components

Pressure and Return SystemComponentsThe force necessary to operate using subsystems isprovided by components of the pressure and returnsystem (Fig 5-8 and 5-9). These components are twoelectrically-driven pumps, pump fuses, pump checkvalves, pressure switch, high-pressure filter, systemrelief valve, pressure indicating components, handpump, accumulators, an accumulator isolation valve,and a return filter.

Electrically-Driven PumpsTwo electrically-driven pumps are installed in theforward end of the left main landing gear wheel well (Fig5-8). The variable-volume pumps are enclosed in ahousing and cover assembly and driven by three-phaseAC motors.

Hand PumpA hand pump is used to pressurize the accumulators andoperate some of the subsystems when theelectrically-driven pumps are not operating (Fig 5-9).

AccumulatorsTwo 400 cubic-inch, piston-type accumulators areinstalled in the service center forward of the reservoir,and are normally charged to approximately 3000 PSI(Fig 5-9).

Case Drain System ComponentsCase drain flow from the electrically driven pumps is

routed back to the reservoir to provide proper cooling

and lubrication for pumps. The case drain system

essentially consist of check valves and a case drain filter

(Fig 5-9).


5-14

SECTION VI ELECTRICAL

SYSTEM

This section consists of information extracted from T.O. 1C-141B-2-24GS-00-1


1. The purpose of the electrical system

2. The function of

• generator drive system

• AC generation system

• DC generation system

• external power system

• emergency generation system

3. The location of the components for

• generator drive system

• AC generation system

• DC generation system

• external power system

• emergency generation system

Electrical System

6-1


6-2

6-1 ELECTRICAL SYSTEM

The electrical system of the C-141 aircraft includes themain generator drive system, the AC and DC generationsystems, the external power system, and the emergency

generation system. The overall functions of thesesystems are described in the following sections.

6-2 GENERATOR DRIVE SYSTEM

A constant speed drive (CSD) unit drives each maingenerator. Their are two interchangeable types of CSDsused on the C-141B. One type is made by GeneralElectric (GE) while the second type is made bySunstrand. Each CSD mounts on an engine accessorydrive case. It is driven at variable speeds but delivers aconstant output speed. CSD operation is monitored on

the flight engineer’s electrical control panel. In case of afailure, the CSD can be electrically disconnected fromthe engine drive. Along with the CSD, the generatordrive subsystem includes the oil supply, an oil tank andpressure regulator, an air-oil cooler, a temperatureregulating pressure bypass valve, oil temperatureindicating, and load controller (Fig 6-1 thru 6-3).

Electrical System

6-3

Fig 6-1, CSD system component locations

1. CSD OIL TANK2. TANK VENT LINE3. CSD OIL PRESSURE REGULATOR4. CSD VENT LINE5. THRUST REVERSER PUMP6. CSD OIL COOLER7. CSD OIL TEMPERATURE

CONTROL8. CSD OIL TEMPERATURE BULB9. CSD10. GENERATOR


6-4

Fig 6-2, CSD oil supply components locations

Electrical System

6-5

Fig 6-3, CSD oil temp ind sys components locations

6-3 AC GENERATION SYSTEM

Three-phase, 200/115-volt, 400-Hz AC power issupplied by four engine-driven generators (Fig 6-5).During ground operations, AC power is supplied fromone of two sources: an auxiliary power unit (APU)driven generator or an external power source (Fig 6-8).During inflight emergencies, AC power is automaticallysupplied from a 2-KVA hydraulically-driven emergencygenerator. The four engine generators, the APU

generator, and external power are controlled from theflight engineer’s electrical control panel (Fig 6-4).Emergency generator operation is normally automatic,but can be controlled from the pilot’s instrument panel.Emergency generator operation is monitored at the flightengineer’s panel. Five voltage regulators, five generatorprotection panels, four load controllers, and a busprotection panel provide control and protection for theAC system(Fig 6-6 and 6-7).


6-6

Fig 6-4, Electrical control panel

Electrical System

6-7

Fig 6-5, Engine driven generator Fig 6-6, Generator load controller locations


6-8

Fig 6-7, Main generator protection panel location

Electrical System

6-9

Fig 6-8, Auxiliary generator location

6-4 DC GENERATION SYSTEM

The normal source of DC power is from twotransformer-rectifier units (Fig 6-9). These units changethree-phase AC voltage to 28 VDC. Each unit is rated at28 volts and 200 amperes. The units also convertexternal or auxiliary generator AC into DC voltageduring ground operation. A battery is provided to startthe APU. The battery is also a standby source for

emergency operation (Fig 6-10). During inflightemergencies. DC power is automatically supplied froma hydraulically-driven emergency generator. The same2-KVA generator also provides emergency AC power.DC power is controlled and monitored at the flightengineer’s electrical control panel.


6-10

Fig 6-9, Transformer rectifier unit location

1. TRANSFORMER RECTIFIER NO 12. TRANSFORMER RECTIFIER NO 2

Electrical System

6-11

Fig 6-10, Battery location

6-5 EXTERNAL POWER SYSTEM

During ground operation, an external power source isnormally used. The external power unit should becapable of supplying 50 KVA three-phase,115/200-VAC, 400-Hz power. Phase sequencing mustmatch that of the aircraft’s generators (A, B, C).Incorrect phase sequence or loss of a phase prevents useof external power on the aircraft. The TRs provide theDC power through normal operation. External power isprovided through a receptacle on the right side of theaircraft (Fig 6-11). Protection circuitry ensures theexternal power is compatible with the aircraft’s

generators. Controls for external power are on the flightengineer’s electrical control panel. External power andauxiliary power cannot be used at the same time. Theexternal power system includes a receptacle, externalpower contactor, part of the bus protection panel, andassociated controls and indicators.

When the external power unit plug is connected to theaircraft, the READY light on the electrical control panelshould come on. The light only turns on after the busprotection panel has checked the phasing and voltage ofthe external power. Placing the power selector switch toEXT connects external power to the main AC tie bus.


6-12

Fig 6-11, External power receptacle

6-6 EMERGENCY GENERATION SYSTEM

The emergency generation system provides AC and DCpower if normal power sources fail. A switch on thepilot’s instrument panel controls manual operation.Warning lights tell the pilot and flight engineer when theemergency generator is on line. The emergency systemis automatically energized due to loss of essential ACbus No. 1 or the emergency DC bus. The emergencysystem can be manually activated during otheremergencies. The emergency system has priority overthe normal power sources to supply the isolated andemergency buses. Emergency power cannot be used tosupply other buses or to charge the battery. The

emergency system is powered by the No. 2 hydraulicsystem. The major components of the system include theemergency generator, hydraulic motor, control solenoidand shutoff valve, frequency sensitive relay, and controlswitches and related circuitry (Fig 6-12).

The emergency power test switch is used with the metersand selector switches to test and monitor the output ofthe emergency generator. The switch has two positions:NORMAL and TEST. Holding the switch to TEST willenergize the emergency generator, but not connect itsoutput to the emergency bus.

Electrical System

6-13

Fig 6-12, Emergency generator location


6-14

SECTION VII OXYGEN SYSTEM

This section consists of information extracted from T.O. 1C-141B-10-35 GS-00-1.


1. The purpose of the oxygen system

2. The function of

• crew oxygen system

• troop oxygen system

• portable oxygen system


• crew oxygen system

• troop oxygen system

• portable oxygen system

Oxygen System

7-1


7-2

7-1 OXYGEN SYSTEM

The oxygen system is used to supply crew members andpassengers with oxygen in case the cabin altitude goesabove a certain limit. Since cabin pressure is notsufficient to force air into the lungs at higher altitudes,the oxygen system is used to supply oxygen at the properflow rates and pressures. There are two independentoxygen systems on the airplane. The crew system is usedexclusively by the flight crew (Fig 7-1) and the troopsystem (Fig 7-4) is used by passengers in the cargo

compartment. Both systems store oxygen in its liquidstate and convert it into a gas before reaching the user(Fig 7-2 and 7-5). Both systems have various indicators,lights, and switches to monitor and test the oxygensystem. External service panels have valves that aid ineasily servicing the oxygen system. Portable oxygenbottles provide mobility for the crew when oxygen isrequired.

Oxygen System

7-3

Fig 7-1, Crew oxygen system components


7-4

TO SYSTEMBUILD-UP

FROM OXYGENCART

12. DRAIN VALVE13. FILL BUILD-UP AND VENT VALVE

Fig 7-2, Crew oxygen system components

7-2 CREW OXYGEN SYSTEM

The crew oxygen system supplies oxygen to the crewmembers. The system is a pressure-demand system,meaning oxygen flows only when the user is demandingoxygen. The system operates at a supply pressure ofapproximately 300 PSI. Oxygen flow and pressure varywith cabin altitude and are controlled at each station by aregulator (Fig 7-1). An indicator and annunciator lightmonitor the quantity of liquid oxygen in a converter (Fig7-3). An indicator test switch tests the indicating system.The normal flow of oxygen under pressure is from the

converter through heat exchangers, where it is warmedto a suitable breathing temperature. From the heatexchangers, the oxygen goes to the nine diluter-demandoxygen regulators. The regulators provide oxygen, ondemand, to the user. The system also has recharger hosesin the flight station, lavatory, and forward cargocompartment for recharging portable oxygen bottles(Fig 7-8).

Oxygen System

7-5

Fig 7-3, Crew LOX system indication

1. OXYGEN QUANTITY LOW ANNUNCIATOR LIGHT2. LIQUID OXYGEN QUANTITY INDICATOR3. PRESS-TO-TEST OXYGEN QUANTITY BUTTON4. MANUALLY OPERATED SHUTOFF VALVE

PILOT’S SIDE CONSOLE

7-3 TROOP OXYGEN SYSTEM

The troop oxygen system, which supplies oxygen topersonnel in the cargo compartment, is a continuousflow distribution system. It operates at a supply pressureof approximately 300 PSI. Distribution pressure varieswith cabin altitude. There are two supply systems andone distribution system (Fig 7-4). Normally, both supplysystems operate during flight. Indicators and lightsmonitor the quantity in the converters. The cargocompartment warning horn and lights indicate that acabin altitude between 12,500 and 14,000 feet has beenreached and oxygen is flowing. Circuits are included totest the quantity indicating system and the warning hornand lights. A separate supply line is provided for

therapeutic oxygen. Normal flow of oxygen is from oneor both converters, under pressure, through heatexchangers. The heat exchangers warm the oxygen to asuitable breathing temperature (Fig 7-5). The oxygencontinues to flow through dual check valves whichinterconnect both supply systems at both continuousflow regulators. The regulators then allow oxygen toflow, either in automatic or manual mode, to thedistribution system. The distribution system is a networkof tubing with spaced outlets to allow installation ofoxygen masks for each seat in the cargo compartment(Fig 7-4).


7-6

Fig 7-4, Troop oxygen distribution system

Oxygen System

7-7

Fig 7-5, Troop oxygen system components


7-8

Fig 7-6, Cargo compartment troop LOX system components

1. NO. 1 LIQUID OXYGEN (LOX) CONVERTERINDICATOR

2. NO. 2 LOX QTY. LOW LIGHT3. OXYGEN ON LIGHT4. NO. 1 LOX QTY. LOW LIGHT5. NO. 1 LOX CONVERTER INDICATOR6. NO. 2 QTY. INDICATOR PUSH-TO-TEST

BUTTON7. OXY LIGHTS AND HORN SWITCH8. HORN SHUTOFF BUTTON9. IND. LIGHTS SWITCH10. NO. 1 QTY. INDICATOR PUSH-TO-TEST

BUTTON11. LOWER REGULATOR12. UPPER REGULATOR13. TROOP OXYGEN CONTROL PANEL14. NO. 1 MANUALLY OPERATED SHUTOFF

VALVE15. NO. 2 MANUALLY OPERATED SHUTOFF

VALVE

7-4 PORTABLE OXYGEN BOTTLES

Five portable oxygen bottles are provided for the crew

(Fig 7-7) and six for the cargo compartment during

emergencies. The bottles can be used for movement

throughout the aircraft when flying at high altitudes. Four

portable oxygen bottles are stowed in the flight station:

two aft of the copilot and two aft of the navigator. One

other bottle is stowed in the lavatory. Each stowage

location has a recharger outlet that is connected to the

crew oxygen system (Fig 7-8).

Oxygen System

7-9

Fig 7-7, Portable oxygen bottles


7-10

Fig 7-8, Oxygen bottle recharger hoses

SECTION VIII AIR CONDITIONING

SYSTEM

This section consists of information extracted from T.O. 1C-141B-2-21GS-00-1.


1. The purpose of the air conditioning system

2. The function of

• bleed air subsystem

• equipment cooling subsystem

• pressurization subsystem

• heating and cooling subsystem

• temperature control subsystem


• bleed air subsystem

• equipment cooling subsystem

• pressurization subsystem

• heating and cooling subsystem

• temperature control subsystem

Air Conditioning System

8-1


8-2

8-1 AIR CONDITIONING SYSTEM

The C-141B air conditioning system includes thefollowing subsystems: equipment cooling, bleed air,pressurization, heating and cooling, and temperaturecontrol (Fig 8-1). The air conditioning system iscontrolled from the flight engineer’s environmental

control panel. Bleed air supplied by the engines, APU, oran external air source is used to operate two airconditioning units, the output of which is mixed withbleed air to heat, cool, ventilate, and pressurize theinterior of the aircraft.


8-3

Fig 8-1, Air conditioning distribution system

8-2 BLEED AIR SYSTEM

Bleed air is hot, compressed air extracted from theengines. It is used to operate the air conditioning unit andmixes with cool air downstream from those units to heatthe flight station and cargo compartment. This air flowinto the aircraft is also used to pressurize the cabin.During ground operations, the APU may be used as asource of bleed air, or a gas turbine compressor may beconnected at the high pressure bleed air connection in theforward end of the left landing gear pod. Bleed air isdistributed from the engines (or from one of the ground

sources) to the various systems by using insulatedstainless steel ducts. The flow of air is controlled byvalves that are operated by positioning switches andcontrols on the flight engineer’s environmental controlpanel.

This system includes engine bleed air shutoff valves,ducts, compensators, an overheat warning system, awing isolation valve, two pressure relief valves, andvarious wing leading edge pressure relief doors.

8-3 EQUIPMENT COOLING SYSTEM

The majority of C-141B avionics and electricalequipment is located in racks below the flight stationfloor. Air from the flight station cools the AC powercontrol panel and flight engineer’s circuit breakerpanels, then is drawn below the flight station to cool theAC electrical load center and underdeck avionicsequipment. Four cooling fans, two for electricalequipment cooling and two for electronic equipmentcooling, move air over and through the equipment (Fig8-2). The two small electrical cooling fans draw in air

from behind the AC power and circuit breaker panelsand discharge it through filters into the AC electricalload center area below deck. The two larger electroniccooling fans draw in air through plenum chambers underthe shelves of the avionics equipment rack. This air isexpelled overboard through a flow control valve. Thissystem includes two electrical cooling fans, twoelectronic cooling fans, a flow control valve, associatedducts, and warning lights on the flight engineer’s panel.


Fig 8-2, Cooling Fans

8-4

4

5

6

1. COOLING FAN FAILURE AVIONICS LIGHT2. COOLING FAN FAILURE ELECTRICAL LIGHT3. ELECTRICAL COOLING FANS4. CHECK VALVES5. DISPOSABLE FILTERS6. ELECTRONIC COOLING FAN

1

2

3

8-4 PRESSURIZATION SYSTEM

The C-141B pressurization system consists primarily ofan automatic controller and a manual controller locatedon the pressurization portion of the flight engineer’senvironmental control panel, and two outflow valvesinstalled in the pressure bulkhead in the aft upper deck(Fig 8-3 and 8-4). The pressurization system maintains

cabin altitude of 8,000 ft. at aircraft altitudes of up to40,000 ft. Pressure is regulated by controlling airflowout of the aircraft through the outflow valves. Thissystem includes an automatic controller, a manualcontroller, two outflow valves, a cabin pressure controlventuri, a control panel, a pressurization control panel,and two negative pressure relief check valves.


8-5

Fig 8-3, Engineer's environmental control panel

1. CABIN PRESS LOW ANNUNCIATORLIGHT

2. CABIN ALT LIMIT OVERRIDE SWITCH3. CLIMB CABIN GAGE4. CABIN ALT GAGE6. CABIN PRES WARNING LIGHT7. CABIN ALT CONTROL8. CABIN PRESSURE CONTROL

INDICATOR9. AUTOMATIC CONTROLLER10. RATE CONTROL11. MANUAL CONTROLLER12. PRESSURIZATION.

CABIN PRESS LOW


8-6

Fig 8-4, Outflow valves

8-5 HEATING AND COOLING SYSTEM

Heating, cooling, and ventilation are provided for theflight station and the cargo compartment on the C-141Bby two parallel air conditioning systems sharing acommon distribution duct system. Normally, the leftsystem and right system are operated in unison, buteither is capable of providing limited heating, cooling,and pressurization to both compartments (Fig 8-5). Onthe ground, pressurized air can be supplied to bothsystems by the APU or by an external air source. Control

of cargo floor temperature is provided by the cargo floorheat system (Fig 8-6 and 8-7). The primary heatexchangers lower the air temperature to approximately

232°C, and refrigerator units incorporating heatexchangers and turbines cool the air further. Condensedmoisture is then extracted from the cold air before aselected amount of warm air is mixed in to yield therequired degree of heating or cooling for the flightstation and cargo compartment.


8-7

Fig 8-5, Left air conditioning pack

1. FAN2. SECONDARY HEAT EXCHANGER3. COOLING TURBINE4. TURBINE BYPASS VALVE


8-8

Fig 8-6, External air connection

Fig 8-7, Cargo floor heat system


8-9

Fig 8-8, Floor heat system components

1. CARGO FLOOR HEAT SHUTOFF VALVE2. FLOOR HEAT MODULATING VALVE3. TEMPERATURE ANTICIPATOR4. FLOOR HEAT EJECTORS5. CONTROL THERMOSTATS6. FLOOR HEAT DUCTS

8-6 TEMPERATURE CONTROL SYSTEM

Various temperature sensors and temperatureanticipators generate signals to temperature controlboxes which, in turn, regulate the temperature of airbeing supplied by the heating and cooling systems. Thetemperature control system also includes the ductingused to distribute conditioned air to the flight station andthe cargo compartment. The percentage of available airdelivered to each station is determined by the diverterand alternate air shutoff valves installed in the left and

right system supply ducts. The flight engineer controlsand monitors the temperature control subsystem withswitches and rheostats on the environmental controlpanel. This system includes two temperature controlboxes and associated sensors, two temperature controlswitches, two temperature control rheostats, a networkof distribution ducts, a diverter valve, an alternate airshutoff valve, and a cargo compartment recirculating airfan (Fig 8-8, 8-9 and 8-10).


8-10

Fig 8-9, Temperature control components


8-11

Fig 8-10, Air conditioning control boxes


8-12

C-141BPROGRAM 5363

AVIONICS SYSTEMS


Scott AFB, IL

PROGRAM 5363AVIONICS SYSTEMS

TABLE OF CONTENTS

SECTION ICOMMUNICATION SYSTEM

PARA TITLE PAGE

1-1 COMMUNICATION SYSTEM . . . . . . 1-3

1-2 HIGH FREQUENCY (HF) SYSTEMS . . 1-3

HF Communication SystemsComponents. . . . . . . . . . . . . . . 1-3

Component Locations. . . . . . . . . . 1-3

HF Communication Systems Antenna . . 1-7

1-3 VERY HIGH FREQUENCY (VHF)SYSTEMS . . . . . . . . . . . . . . . . . 1-7

Components. . . . . . . . . . . . . . .1-10


1-4 SECURE VOICE (KY-58 SYSTEM) . . 1-10

Components. . . . . . . . . . . . . . .1-10

Component Location. . . . . . . . . . 1-10

1-5 MULTIPLE RADIO TRANSMITSYSTEM . . . . . . . . . . . . . . . . . 1-10

1-6 ULTRA HIGH FREQUENCY (UHF)SYSTEMS . . . . . . . . . . . . . . . . 1-10

Components. . . . . . . . . . . . . . .1-11


1-7 INTERPHONE SYSTEMS . . . . . . . . 1-12

Components. . . . . . . . . . . . . . .1-13


1-8 PUBLIC ADDRESS SYSTEM . . . . . . 1-14

Components. . . . . . . . . . . . . . .1-14


1-9 STATIC DISCHARGING SYSTEM. . . 1-15

Components. . . . . . . . . . . . . . .1-15


1-10 UHF SATELLITE TERMINAL SYSTEM(USTS) ANTENNA SYSTEM . . . . . . 1-16

Components. . . . . . . . . . . . . . .1-16


SECTION IINAVIGATION SYSTEM

PARA TITLE PAGE

2-1 NAVIGATION SYSTEM . . . . . . . . . 2-3

2-2 PITOT-STATIC SYSTEM. . . . . . . . . 2-3

2-3 STANDARD CENTRAL AIR DATACOMPUTER (SCADC) SYSTEM. . . . . 2-6

2-4 TOTAL AIR TEMPERATURESYSTEM . . . . . . . . . . . . . . . . . 2-10

2-5 RADAR ALTIMETER SYSTEM . . . . 2-11

Receiver-Transmitter . . . . . . . . . . . 2-11

Radar Altimeter Indicator . . . . . . . . 2-11

iii

Signal Data Converter . . . . . . . . . . 2-11

2-6 ATTITUDE AND HEADINGREFERENCE SYSTEM (AHRS) . . . . 2-13

Displacement Gyroscope Assembly . . 2-13

Compass System Controller. . . . . . 2-13

AHRS Power Control Panel. . . . . . 2-13

Magnetic Azimuth Detector. . . . . . 2-13

Remote Magnetic Compensator. . . . 2-13

2-7 FLIGHT DIRECTOR SYSTEM . . . . . 2-16

Navigation Selector Panels. . . . . . . 2-16

Rate Gyroscope Transmitter. . . . . . 2-16

Power Adequacy Indicators. . . . . . . 2-16

Test Switch. . . . . . . . . . . . . . .2-16

Slaving Switch . . . . . . . . . . . . . 2-16

Elevator Position Transmitter. . . . . . 2-16

Flight Director Computer . . . . . . . 2-16

Horizontal Situation Indicator (HSI) . . 2-16

Attitude Direction Indicator (ADI). . . 2-16

2-8 GLIDE SLOPE SYSTEM . . . . . . . . 2-20

Glide Slope Receiver. . . . . . . . . . 2-20

Glide Slop Antenna. . . . . . . . . . . 2-20

2-9 MARKER BEACON SYSTEM(MKR BCN) . . . . . . . . . . . . . . . 2-22

Marker Beacon Receiver. . . . . . . . 2-22

Marker Beacon Antenna. . . . . . . . 2-22

Marker Beacon Indicators. . . . . . . 2-22

MKR BCN HIGH/LOW Switch . . . . 2-22

2-10 IDENTIFICATION FRIEND ORFOE (IFF) . . . . . . . . . . . . . . . . . 2-24

IFF Control Panel. . . . . . . . . . . . 2-24

IFF Receiver-Transmitter. . . . . . . . 2-24

IFF Transponder Test Set. . . . . . . . 2-24

IFF Antennas. . . . . . . . . . . . . .2-24

IFF Antenna Switching Unit. . . . . . 2-24

Mode 4 Computer. . . . . . . . . . . . 2-24

2-11 APS-133 COLOR WEATHER RADARSYSTEM . . . . . . . . . . . . . . . . . 2-28

Receiver-Transmitter. . . . . . . . . . 2-28

Multifunction Display . . . . . . . . . 2-28

Antenna. . . . . . . . . . . . . . . . .2-28

Radar Control Panel. . . . . . . . . . 2-28

2-12 GROUND PROXIMITY WARNINGSYSTEM (GPWS) . . . . . . . . . . . . 2-32

2-13 INERTIAL NAVIGATION SYSTEM(INS) . . . . . . . . . . . . . . . . . . . 2-33

Inertial Navigation SystemComponents. . . . . . . . . . . . . . .2-33

2-14 AUTOMATIC DIRECTION FINDERSYSTEM (ADF) . . . . . . . . . . . . . 2-38

2-15 VHF NAVIGATION SYSTEM . . . . . 2-41

VHF Navigation Receivers. . . . . . . 2-41

VHF COMM/NAV Control Panels. . . 2-41

VHF NAV VOR Tail Antennas . . . . 2-41

VHF NAV LOC Nose Antenna. . . . . 2-41

VHF NAV Antenna Switch. . . . . . . 2-41

2-16 AN/ARN-118 TACTICAL AIRCRAFTNAVIGATION SYSTEM(TACAN SYSTEM) . . . . . . . . . . . 2-45

2-17 AN/APN 169C INTRAFORMATIONPOSITIONING SYSTEM (SKE). . . . . 2-47

Primary Control. . . . . . . . . . . . . 2-47

Secondary Control. . . . . . . . . . . 2-47

iv

Flight Command Indicator (FCI). . . . 2-47

Flight Command Repeater (FCR). . . 2-47

Relative Range Indicator (RRI). . . . . 2-47

Audio Amplifier . . . . . . . . . . . . 2-47

Auxiliary Interface Unit (AIU) . . . . . 2-47

Receiver-Transmitter. . . . . . . . . . 2-52

Coder-Decoder. . . . . . . . . . . . . 2-52

Electrical Equipment Rack. . . . . . . 2-52

Antenna. . . . . . . . . . . . . . . . .2-52

Antenna Pedestal. . . . . . . . . . . . 2-52

2-18 FUEL SAVINGS ADVISORY SYSTEM(FSAS) . . . . . . . . . . . . . . . . . . 2-53

SECTION IIIINDICATING AND RECORDING SYSTEM

PARA TITLE PAGE

3-1 INDICATING AND RECORDINGSYSTEM . . . . . . . . . . . . . . . . . . 3-3

3-2 COCKPIT VOICE RECORDER (CVR)SYSTEM . . . . . . . . . . . . . . . . . . 3-3

Components. . . . . . . . . . . . . . . 3-3

Cockpit Voice Recorder (CVR). . . . . 3-3

Cockpit Voice Recorder (CVR)Control Unit . . . . . . . . . . . . . . . 3-3

Cockpit Voice Recorder TESTSwitch . . . . . . . . . . . . . . . . . . 3-3

3-3 DIGITAL FLIGHT DATA RECORDER(DFDR) SYSTEM . . . . . . . . . . . . . 3-6

3-4 TAKEOFF WARNING SYSTEM . . . . . 3-6

3-5 LIFE HISTORY RECORDER SYSTEM(LHRS) . . . . . . . . . . . . . . . . . . . 3-6

SECTION IVAUTOMATIC FLIGHT CONTROL SYSTEM

PARA TITLE PAGE

4-1 AUTOMATIC FLIGHT CONTROLSYSTEM . . . . . . . . . . . . . . . . . . 4-3

4-2 AUTOPILOT. . . . . . . . . . . . . . . . 4-3

Autopilot Controls and Indicators. . . . 4-3

AFCS Control Panel Procedures. . . . . 4-3

Control Wheel Steering. . . . . . . . . 4-3

Autopilot Mode Selection. . . . . . . . 4-3

Autopilot Operational Modes. . . . . . 4-3

Heading Hold Mode. . . . . . . . . . . 4-3

Altitude Hold Mode . . . . . . . . . . . 4-3

Mach Hold Mode. . . . . . . . . . . . . 4-3

Pitch Off Mode. . . . . . . . . . . . . . 4-3

Lateral Off Mode . . . . . . . . . . . . 4-4

Navigation Select Operation. . . . . . . 4-4

Heading Select Mode. . . . . . . . . . 4-4

(VOR)(ILS) Mode . . . . . . . . . . . . 4-4

INS Select Mode. . . . . . . . . . . . . 4-4

Fuel Savings Advisory SystemMode (FSAS). . . . . . . . . . . . . . . 4-4

4-3 YAW DAMPER SYSTEM . . . . . . . . 4-6

Theory of Operation. . . . . . . . . . . 4-6

Warning System. . . . . . . . . . . . . 4-6

Testing . . . . . . . . . . . . . . . . . . 4-6

4-4 AUTOMATIC THROTTLE SYSTEM(ATS) . . . . . . . . . . . . . . . . . . . 4-6

Airspeed Trim Assembly. . . . . . . . . 4-6

ATS Computer/Amplifier Unit. . . . . . 4-6

v

Motor Generator Assembly. . . . . . . 4-6

Clutch Pack Assembly. . . . . . . . . . 4-6

4-5 FUEL SAVINGS ADVISORYSYSTEM (FSAS) . . . . . . . . . . . . . 4-9

FSAS INS Navigation Aiding. . . . . . 4-9

FSAS Told Card Information. . . . . . 4-9

Flight Planning. . . . . . . . . . . . . . 4-9

Flight Profiles . . . . . . . . . . . . . . 4-9

FSAS INS Radar Displays. . . . . . . . 4-9

FSAS Components. . . . . . . . . . . . 4-9

FSAS Computer. . . . . . . . . . . . . 4-9

FSAS Display Interface Unit. . . . . . . 4-9

FSAS Display Interface Control Unit . . 4-9

FSAS Control Display Unit. . . . . . . 4-10

4-6 ALL WEATHER LANDING SYSTEMS(AWLS) . . . . . . . . . . . . . . . . . . 4-10

AWLS Components. . . . . . . . . . . 4-10

AWLS Test Programmer and LogicComputer (TPLC). . . . . . . . . . . . 4-10

Rotation-Go-Around Computer(RGA). . . . . . . . . . . . . . . . . .4-10

Flare Computer. . . . . . . . . . . . . 4-10

vi

SECTION I COMMUNICATION

SYSTEM

This section consists of information extracted from T.O. 1C-141B-23GS-00-1.


1. The purpose of the communication system

2. The function of

• HF communication system

• VHF communication

• secure voice (VHF,UHF) KY-58 system

• UHF communication system

• interphone system

• public address

• static discharger

• UHF satellite terminal system (USTS)


• HF communication system

• VHF communication

• secure voice (VHF,UHF) KY-58 system

• UHF communication system

• interphone system

• public address

• static discharger

• UHF satellite terminal system (USTS)

Communication System

1-1


1-2

1-1 COMMUNICATION SYSTEM

The C-141B communication system provides air-to-air,air-to-ground and intra-aircraft communications, securevoice, and multiple radio transmit. A more detaileddescription of each system is presented in subsequentsections of this program. System and type identificationfor installed communication equipment are as follows:

SYSTEM . . . . . . . . . . . . . TYPE

HF COMM. . . . . . . . . . . HF-102

VHF COMM . . . . . . . . . . ARC-186 (V)

SECURE VOICE(VHF, UHF) . . . . . . . . . . KY-58

MULTIPLE RADIO TRANSMIT

UHF COMM. . . . . . . . . . ARC-164 (V)

INTERPHONE. . . . . . . . . AIC-18A

PUBLIC ADDRESS. . . . . . AIC-13

STATIC DISCHARGING

UHF SATELLITE TERMINAL SYSTEM (USTS)

SECURE VOICE (HF). . . . . KY-75A

1-2 HIGH FREQUENCY (HF) SYSTEMS

The No. 1 and No. 2 HF systems provide long rangeradio-telephone communication. Both systems operateon either conventional amplitude modulation (AM) orsingle side band (SSB) modes. SSB mode concentratesincreased power to the side band while suppressing thecarrier. SSB is used to communicate with stations havingSSB capability. AM mode provides less side bandpower, but enables communication with stations nothaving SSB capability. Anyone of 28,000 frequencychannels can be selected from the HF control panels. HFNo. 1 and No. 2 are provided a secure voice system. Thesystem encrypts and decodes classified transmissionsaircrew and external transmitter/receivers. On aircraftequipped with multiple radio transmit capability, HF No.l can transmit and receive simultaneously with VHF andUHF No. l systems.

HF Communication SystemsComponentsThe following major components comprise the HFcommunication system:

COMPONENT. . . . . . . . . . QUANTITY

Control Panel. . . . . . . . . . 2

Accessory Unit. . . . . . . . . 2

Antenna . . . . . . . . . . . . 1

Antenna Coupler. . . . . . . . 2

Lightning Arrestor Relay Unit . 1

Transceiver. . . . . . . . . . . 2

KY-75 Secure Voice. . . . . . 4

KY-75 Processor. . . . . . . . 2

Remote Relay Box. . . . . . . 2

Component LocationsHF No. l and No. 2 communication systems controlpanels are installed in the center console (Fig 1-1). HFNo. 1 secure voice remote control unit (KY-75 RCU) isinstalled in the pilot’s side console. HF No. 2 KY-75RCU is installed in the copilot’s side console. Dual HFNo. l and No. 2 KY-75 RCU panels are installed in thenavigator’s instrument panel. HF No. l and No. 2 KY-75processors are mounted under the navigator’s table. HFNo. l and No. 2 receiver transmitter are installed on theright avionics equipment rack in the underdeck area (Fig1-3). HF No. l and No. 2 RE-978 remote relay boxes areinstalled on the left avionics equipment rack in theunderdeck area. HF No. l KY-75 accessory unit isinstalled at the top LH side of FS 1600. No. 2 accessoryunit is mounted on the RH side at FS 1600. HF No. l andthe No. 2 antenna couplers are installed on the antennalightning arrestor relay unit. Both antenna couplers andlightning arrestor relay unit are located in the T-tail,center and aft of the HF antenna and bullet location (Fig1-2).


1-3


1-4

Fig 1-1, HF radio control panel location


1-5

Fig 1-2, HF system components


1-6

Fig 1-3, HF receiver transmitters

HF Communication Systems AntennaBoth HF No. 1 and No. 2 share a common probe antennaextending forward from the top of the T-tail. Both HF

systems can receive simultaneously, but relay action inthe RT mount and left-side coupler mount/relay unitallows only one HF system to transmit at a time.

1-3 VERY HIGH FREQUENCY (VHF) SYSTEMS

The VHF communication systems are used to transmitand receive amplitude modulated (AM) radiocommunications. Voice (audio) signals to be transmittedare supplied through the interphone system. Audioreceived from other stations is supplied to the interphonesystem headsets. Two complete VHF communicationsystems (No. l and No. 2) are installed. Both systemsoperate independently and can be used simultaneously.Each system provides channels spaced 25 KHz apart,from 116.00 to 151.975 MHz. The transmitter poweroutput is approximately 10 watts nominal of RF energy.VHF No. 1 or No. 2 may be operated by the pilot (No. l)or copilot (No. 2) from control panels on the centerconsole. VHF communication is used primarily during

landing, takeoff, air-to-air, and air-to-groundoperations. Because of the straight-line radiationcharacteristics of VHF radio waves, communication isrestricted to line-of-sight distances. Averagecommunication distances from aircraft-to-ground areapproximately 40 miles at 1,000 feet to 135 miles at10,000 feet altitude. VHF No. l and No. 2 are provided aKY-58 secure voice system. The system encrypts anddecodes classified transmissions between aircrew andexternal transmitter/receivers. VHF and UHF radioKY-58 secure voice systems are discussed with multipleradio transmit capability. VHF No. 1 can transmit andreceive simultaneously with HF and UHF No. 1 systems.


1-7

Fig 1-4, VHF system control panels

1. VHF CONTROL PANEL No. 12. VHF CONTROL PANEL No. 23. NAV SLEW SWITCH ACTIVE/INACTIVE INDICATOR4. COMM SLEW SWITCH ACTIVE/INACTIVE INDICATOR5. NAV VOLUME CONTROL6. SLEW ENABLE SELECT SWITCH7. DISPLAY BRIGHTNESS CONTROL8. MODE SELECTOR SWITCH9. HUNDREDTHS/THOUSANDTHS PRESET

CHANNEL SLEW SWITCH10. TENTHS SLEW SWITCH11. MEMORY LOAD SWITCH12. UNITS SLEW SWITCH13. TENTHS SLEW SWITCH14. COMM MODE CONTROL SWITCH15. SQUELCH DISABLE/TONE SELECT SWITCH

16. COMM VOLUME CONTROL17. NAV\LAMP TEST SWITCH18. NAV DISPLAY19. COMM DISPLAY


1-8

Fig 1-5, VHF components

Fig 1-6, VHF antenna


1-9

Fig 1-7, KY-58 secure voice component location

ComponentsThe following major components comprise the VHFcommunications system:


Transceiver. . . . . . . . . . . 2

Control Panel. . . . . . . . . . 2

Antenna . . . . . . . . . . . . 2

KY-58 Secure Voice RemoteControl Unit (KY-58 RCU) . . 4

KY-58 Local/RemoteSwitch . . . . . . . . . . . . . 1

KY-58 Processor. . . . . . . . 4

Component LocationsVHF No. 1 and No. 2 communication systems controlpanels are installed in the center console (Fig 1-4).Transceivers for both systems are installed on the left,center avionics equipment rack in the underdeck area.No. l VHF system antenna is mounted under the fuselagebetween the main landing gear pods at FS 924. No. 2VHF antenna is mounted on top of the centerwing area atFS 882 (Figs 1-5 and 1-6).

1-4 SECURE VOICE (KY-58 SYSTEM)

The Ky-58 secure voice system encrypts and decodestransmissions through the VHF and UHF transceivers.Crypto variables are loaded into the KY-58 processorsfrom a KOI-18 tape reader or a KYK-13 transfer device.Fill connectors are provided on each processor controlpanel for connecting the loading units. The KY-58system has two basic modes of operation: plain or cipher.The plain mode is used during normal communication.The cipher mode is used when secure voicecommunications are required. The receiving stationmust be properly equipped to receive transmissions inthe cipher mode, communications are encrypted andcannot be understood without KY-58 to decode. Incipher, transmitted voice is always encrypted andreceived voice is decrypted so it can be understood. Anyplain text voice being received will also be heard. Whencipher only mode is used, plain text messages beingreceived will not be heard.

ComponentsThe following major components comprise the UHF andVHF secure voice system:


VHF KY-58 Processor. . . . . 2

UHF KY-58 Processor. . . . . 2

KY-58 RCU . . . . . . . . . . 4

LOCAL/REMOTE Switch. . . 1

Component LocationA KY-58 processor and remote control unit (RCU) areprovided for each VHF and UHF transceiver. Theprocessors are located in the navigator’s panel. A bypassunit is mounted adjacent to each processor behind thenavigator’s panel. Two RCU’s are mounted in each ofthe pi lot’s and copi lot’s s ide consoles . ALOCAL/REMOTE switch is located in the navigator’spanel above the VHF No. 1 KY-58 processor (Fig 1-7).

1-5 MULTIPLE RADIO TRANSMIT SYSTEM

This system has been deactivated.

1-6 ULTRA HIGH FREQUENCY (UHF) SYSTEMS

Dual UHF radio systems are used for line-of-sight,radio-telephone communication. The systems aredesignated UHF No. 1 and UHF No. 2. Communicationrange varies with altitude. As the altitude of the aircraftincreases, the line-of-sight range increases. The systems

provide two-way, AM (amplitude modulation),double-side-band transmission and reception. UHFradio system Have Quick capability provides air-to-airand air-ground-air, jam-resistant voice communication.The Have Quick UHF radio system has a frequency


1-10

hopping capability. Frequency hopping is a techniquewhere the frequency being used for communication israpidly changed many times per second. Frequencyhopping is implemented by storing a pattern offrequencies for a given day within every Have QuickUHF radio and utilizing this pattern according to the timeof day. The frequency used at a particular instantdepends on the precise time of day; both UHF radios of aHave Quick communicat ion l ink must havesynchronized clocks. Universal Coordinated Time(UCT) has been adopted for Have Quick UHF radiotime. A secure voice system encrypts and decodesclassified transmissions between pilot, copilot ornavigator, and external transmitter/receivers.

ComponentsThe following major components comprise the UHFcommunication systems:


Transceiver. . . . . . . . . . . 1

Antenna . . . . . . . . . . . . 1

Component LocationsUHF No. 1 and No. 2 communication systemstransceivers are installed in the center console (Fig 1-8).No. 1 UHF antenna is mounted under the fuselage, aft ofthe main landing gear pods at FS 1025. No. 2 UHFantenna is mounted on top of the center wing area at FS670 (Fig 1-9). A multichannel (conferencing) capability


1-11

Fig 1-8, UHF system control panels

1. No. 1 UHF RADIO2. No. 2 UHF RADIO

Fig 1-9, UHF system antenna location

1. UHF RADIO ANTENNA No. 12. UHF RADIO ANTENNA No. 2

permits a receiver to accept multiple transmissions. Thisis implemented in the transmitter where it is recognizedif the net is already in use. If so, the transmitter sidesteps

by one 25 KHz channel. Since the receiver has a wideband mode, the second signal is received in addition tothe original.

1-7 INTERPHONE SYSTEMS

The interphone system provides communication amongthe crew members in the flight station and cargocompartment. Communication with the ground crew isalso possible through external receptacles (Fig 1-10). Inaddition to interphone communication, the systemprovides switching and mixing facilities for selectingand monitoring the radio communication and navigationsystems. A loudspeaker is used in the flight station tomonitor the pilot’s audio signals. The speaker also servesas an audible warning of an engine fire, aircraftoverspeed, or under spoiler speed. There are ten

interphone stations within the aircraft. Five stations arein the flight station area, three in the cargo area, one inthe avionics equipment rack, and one in the verticalstabilizer. Three external receptacles are installed forground maintenance. One is on the lower left side of thefuselage just forward of the nose wheel, one is near theforward end of the main wheel well, and one is in theright main wheel well. The aft receptacles connect intothe cargo area panels and the forward receptacleconnects into the avionics equipment rack. The flight


1-12

Fig 1-10, Interphone connections

1. EXTERNAL RECEPTACLE2. COPILOT’S CONTROL PANEL3. FLIGHT ENGINEER’S CONTROL PANEL4. EXTERNAL CONNECTION FOR UARRSI TESTER

USE5. RIGHT JUMPMASTER’S\LOADMASTER’S

CONTROL PANEL6. VERTICAL STABILIZER AUXILIARY CONTROL

PANEL7. LEFT JUMPMASTER’S\LOADMASTER

CONTROL PANEL WITH NURSE RECEPTACLE8. FORWARD CREW DOOR CONTROL PANEL9. AVIONICS COMPARTMENT CONTROL PANEL

10. NAVIGATOR’S AND OBSERVER’S CONTROLPANEL

11. PILOT’S CONTROL PANEL

station loud speaker is installed overhead in the flightstation aft of the pilot’s position.

ComponentsThe following major components comprise theinterphone communication system (Fig 1-11):

COMPONENTS . . . . . . . . . QUANTITY

Control Panel. . . . . . . . . . 9

Monitor Panel . . . . . . . . . 5

Auxiliary Control Panel. . . . 1

Control Wheel Switch. . . . . 2

Nose Steering Wheel Switch. . 1

Foot Switch . . . . . . . . . . 2

Hand-held Microphone. . . . 5

Loudspeaker. . . . . . . . . . 1

Component LocationsThe flight station interphone positions are pilot’s,copilot’s, navigator’s, flight engineer’s, and observer’s.The flight engineer’s and navigator’s stations eachcontain a second headset and microphone jack for use byinstructors. Each station is provided with an interphonecontrol panel and monitor panel, except in the flightengineer’s station. The flight engineer has a controlpanel only. Cargo area stations are provided controlpanels. The left jumpmaster’s/loadmaster’s panelincludes an interphone cord receptacle to provide a flightnurse walk around interphone capability. The avionicsequipment rack contains a control panel and monitorpanel. An auxiliary control panel is in the verticalstabilizer. All of the monitor panels are identical. Thecontrol panels are the same internally, but the panelmarkings vary between the flight station and cargo areastations. The primary difference between the interphonestations is in the aircraft wiring. The avionics equipmentrack panels provide the same operation as those in theflight station.


1-13

Fig 1-11 Interphone components location

1. NOSE WHEEL STEERINGINTERPHONE BUTTON

2. CONTROL WHEEL SWITCH3. HAND-HELD MICROPHONE

1-8 PUBLIC ADDRESS SYSTEM

The public address system provides one-waycommunication in the cargo area through sixloudspeakers located in the cargo compartment (Fig1-12) The main panel is located on the flight engineer’sinstrument panel. The aft speakers may be used tocommunicate with the areas to the rear of the aircraft.Auxiliary panels are located at the forward crewentrance door and the lef t and rightjumpmaster/ loadmaster stations. Microphoneconnections to the system are made through theinterphone system.

ComponentsThe following major components comprise the publicaddress system:


Main Control Panel . . . . . . 1

Auxiliary Control Panel. . . . 3

Power Amplifier . . . . . . . . 3

Loudspeaker. . . . . . . . . . 6

Component LocationsThe main public address (PA) control panel is installedon the flight engineer’s instrument panel. Auxiliarycontrol panels are located at the cargo compartmentinterphone stations. PA system amplifiers are mountedon the LH avionics equipment rack. Loudspeakerselector relays are installed on the interphone junctionbox located in the RH avionics compartment.Loudspeakers are located in the cargo compartment. Apublic address control switch is in the pilot’s sideconsole.


1-14

Fig 1-12, PA system component location

1-9 STATIC DISCHARGING SYSTEM

Thirty-five static dischargers are located at variouspoints along the trailing edges of the wings andempennage. The static dischargers help prevent radiointerference by dissipating accumulated static electricityinto the atmosphere.

ComponentsEach static discharger consists of a retainer and a tipassembly. The retainer is bonded to the aircraft surfaceby means of an electrically conductive adhesive toprovide a low resistance bond between the staticdischarger and the aircraft surface. The 35 retainers areriveted to the aircraft surface in addition to being

bonded. The tip assembly of each static discharger issecured in the retainer by means of a setscrew.

Component LocationsThe static dischargers are located in the aircraft asfollows: five at the outboard ends of the right and leftelevator trailing edges, two at the tips of the right and leftends of the horizontal stabilizer, two near the bottom ofthe rudder’s trailing edge, six at the outboard trailingedges of the right and left wings, three at the tips of theleft and right wings, and one at the aft end of theempennage bullet (Fig 1-13).


1-15

Fig 1-13, Static discharging system

1-10 UHF SATELLITE TERMINAL SYSTEM (USTS)

ANTENNA SYSTEM

The UHF satellite terminal systems (USTS) antennasubsystem links the aircraft’s UHF SATCOMtransceiver with the SATCOM antenna (Fig 1-14). AllC-141B model aircraft are equipped with SATCOMantenna. Special OPS personnel install a portable radioto complete the system.

ComponentsThe following major components comprise theaircraft-installed USTS antenna system:


Antenna DMC34-8-1. . . . . . 1

Antenna DMC34-3-b. . . . . . 1

USTS Interface Panel. . . . . 1

Amplifier Equipment Shelf . . 1

RF Amplifier AM-7388/A. . . 1

Quadrature DMH 26-1. . . . . 1

Dummy Load 376NF . . . . . 1

Component LocationsThe externally mounted UHF antenna has three cablesconnected to a feed-through doubler on the aircraft’soverhead structure. The amplifier equipment shelf ismounted on the aircraft frame below the antenna at FS698. The USTS Interface Panel is mounted on the rightcargo compartment wall between FS458E and FS478E.


1-16

Fig 1-14, UHF Satellite terminal system

SECTION II NAVIGATION

SYSTEM

This section consist of information extracted from T.O. 1C-141B-34GS-00-1.


1. The purpose of the navigation system

2. The function of

• pitot-static system

• central air data computer

• total air temperature

• radar altimeter system

• attitude and heading reference system

• flight director system

• glide slope system

• marker beacon system

• identification friend or foe system

• color weather radar system

• ground proximity warning system

• inertial navigation system

• automatic direction finder system

• VHF navigation system

• tactical air navigation (TACAN)

• AN\APN169C intraformation positioning system

• fuel savings advisory system


• total air temperature

• radar altimeter system

• attitude and heading reference system

• flight director system

• glide slope system

• marker beacon system

• identification friend or foe system

• color weather radar system

• VHF navigation system

• AN\APN169C intraformation positioning system

Navigation System

2-1


2-2

2-1 NAVIGATION SYSTEM

These various systems collectively provide informationand signals for flying the aircraft to any place on earthwith optimum fuel economy and in optimum time. Theoverall functions and operations of the navigationsystem are described in this section. A more detaileddescription of each system is presented in subsequentsections of this program.

Pitot-Static

Standard Central Air Data Computer (SCADC)

Total Air Temperature

Radar Altimeter

Attitude and Heading Reference System (AHRS)

Flight Director

Glide slope (G/S)

Marker Beacon

IFF

Weather Radar

Ground Proximity Warning System (GPWS)

Inertial Navigation System (INS)

Automatic Direction Finder (ADF)

VHF Navigation

TACAN

Intraformation Positioning

Fuel Savings Advisory System (FSAS)

2-2 PITOT-STATIC SYSTEM

The pitot-static system provides pitot and static airpressure for the pilot’s and copilot’s central air datacomputer and other functional units. The externalpressure is picked up by four pitot-static tubes, just aft ofthe crew’s entrance door (Fig 2-1). The tube heads andmasts are provided with heater elements for deicing thesystem. The pitot-static system provides a source ofexternal total (pitot) air pressure and external static(atmospheric) air pressure. The pitot-static systemroutes the air pressure to indicators and equipmentlocated inside the aircraft. The air pressure is convertedinto airspeed and altitude information. The pitot-static

system provides both types of air pressure to the standbyairspeed indicator, SCADC, and an altitude airspeedtransducer. Static pressure is supplied to the flightengineer’s and navigator’s altimeters and a cabindifferential pressure transducer. Two pitot-staticsystems are installed on the aircraft: one is utilized toprovide data for use by the pilot and the other providesdata for use by the copilot.

The pitot-static system consists of pitot-static probes,system plumbing (Fig 2-2), drainage provisions, shutoffvalves, and anti-icing provisions.

Navigation System

2-3


2-4

Fig 2-1, Pitot-static indicators

Navigation System

2-5

Fig 2-2, Pitot-static system plumbing

2-3 STANDARD CENTRAL AIR DATA COMPUTER (SCADC)

SYSTEM

The SCADC supplies primary flight information to thepilot’s and copilot’s vertical scale flight instruments(VSFI), as well as control signals to other systems suchas the automatic flight control system (AFCS). Thepilot’s and copilot’s VSFI indicate aircraft speed interms of both mach number and airspeed, barometricaltitude, and the rate at which aircraft altitude change.Two complete and independent SCADC systems areinstalled in the aircraft (Fig 2-3, 1 of 4 thru 4 of 4). The

central air data system is active whenever power isavailable. The system derives all of its output parametersfrom inputs of pitot-static air pressures and airtemperature. The SCADC computes indicated airspeed,true airspeed by adding air temperature corrections tothe indicated airspeed calculations, barometric altitude,and altitude rate. The SCADC is an all digital-electronicLRU and has no electromechanical parts.


2-6

Fig 2-3, (1 of 4), SCADC components and locations

Navigation System

2-7



2-8


Navigation System

2-9


2-4 TOTAL AIR TEMPERATURE SYSTEM

There are two total air temperature indicators one on thepilot’s instrument panel and one on the flight engineer’sinstrument panel (Fig 2-4). Total air temperature signalsare provided to SCADC systems and to the fuel savingsadvisory system (FSAS). The total air temperature is

served by resistive elements in two probes, one justforward of the crew entrance door on the left side of theaircraft. The other probe is in a similar position on theright side. Each probe has a resistive heating element foranti-icing.


2-10

Fig 2-4, Total air temperature components

2-5 RADAR ALTIMETER SYSTEM

The combined altitude radar altimeter system (CARA)AN/APN-232(V) measures the aircraft’s altitude abovethe ground. The altimeter is an altitude tracking andindicat ing radar . Alt i tude is measured by areceiver-transmitter and is visually displayed by anindicator on the pilot’s center instrument panel. Theindicator displays altitudes between zero and 50,000feet. The system is set to show zero feet at touchdown.Radar altimeter operations are not affected byatmospheric or barometric conditions. Altitude datasignals are sent to the ground proximity warning system(GPWS) and the all weather landing system (AWLS).

The radar altimeter system consists of a radar altimeterreceiver-transmitter, a radar altimeter height indicator, asignal data converter, and a pair of antennas .

Receiver-Transmitter

The receiver-transmitter is located on the left side of thecargo compartment at FS 1038 and contains the

functional circuitry that generates the transmittedsignals and processes the received signals to achievealtitude measurement.

Radar Altimeter Indicator

The radar altimeter indicator is on the pilot’s centerinstrument panel. The indicator has a set control knobthat enables the system when the knob is rotatedclockwise to on.

Signal Data Converter

The signal data converter located on the left side of thecargo compartment at FS 1038, contains the functionalcircuitry that converts the self test and altitude outputs ofthe receiver-transmitter into the radar altitude signalsutilized by the AWLS and tracking circuits to providerange (altitude) information.

Navigation System

2-11


2-12

Fig 2-5, Altimeter components and locations

2-6 ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS)

AHRS is the sole source of magnetic heading for the autopilot and flight recorder system. The magnetic headingsignal is also backup to the INS No. 1 and No. 2 (forattitude and heading), the horizontal situation indicators(HSI), the bearing distance heading indicators (BDHI),the tactical air navigation (TACAN) system, the VHFomni range (VOR) navigation system, and the allweather landing system (AWLS). Additionally, pitchand roll attitude signals are provided to the attitudedirector indicators (ADI), the test programmer logiccomputer (TPLC), and the weather radar system. TheAHRS also provides a turn signal that is used by thepilot’s ADI rate of turn indicator.

The attitude and heading reference system (AHRS)provides a backup source of magnetic heading(compass) and attitude signals to other aircraft systems.Magnetic heading signal is used for aircraft navigation.The AHRS processes signals from a detector that sensesthe earth’s magnetic field. The AHRS also has adisplacement gyroscope all attitude platform whichprovides pitch, roll, and azimuth signals. The AHRS canalso provide unstabilized magnetic heading, but this is anabnormal or emergency mode of operation. The systemmonitors its own operation and provides a warningdiscrete if output information is unreliable.

The AHRS consists of the displacement gyroscopeassembly, compass system controller, power controlpanel, a magnetic azimuth detector, and a remotemagnetic compensator (Fig 2-6, 1 of 2 thru 2 of 2).

Displacement Gyroscope Assembly

The displacement, gyroscope, electronic controlamplifier, and the mount are packaged together as thedisplacement gyroscope assembly located in the centeravionics equipment rack (Fig 2-6, 1 of 2).

Compass System Controller

The CSC provides all of the operating controls andindicators for AHRS. Located on the copilot’s sideconsole, the controls consist of two toggle switches (N/Shemishere switch, Mag VAR switch), a rotary switch(mode switch), a potentiometer control (latitudecontrol), and a sync/heading set control (-HGD+/Push toSYNC control) (Fig 2-6, 2 of 2). One indicator meter ison the panel (SYN indicator).

AHRS Power Control Panel

The AHRS power control panel is located on thecopilot’s side console. It has two toggle switches: PWRand FAST-ERECT/NORM. The PWR switch is a twoposition ON-OFF switch that supplies primary power tothe AHRS.

Magnetic Azimuth Detector

The magnetic azimuth detector (MAD) is a device thatsenses the earth’s magnetic flux. Pendulums mounted tostay in a horizontal position, the MAD is located in theleft wing tip to isolate it as much as possible frommetallic structure (Fig 2-6, 1 of 2).

Remote Magnetic Compensator

The remote magnetic compensator provides ,adjustment for index error during compass swing. It islocated in the left underdeck avionics compartment (Fig2-6, 1 of 2).

Navigation System

2-13


2-14

Fig 2-6, (1 of 2), AHRS components and locations

Navigation System

2-15

Fig 2-6, (2 of 2), AHRS components and locations

2-7 FLIGHT DIRECTOR SYSTEM

The flight director system provides the pilot and copilotwith a pictorial display of data required to performinstrument flight maneuvers. Two completely separateand independent flight director systems exist on theaircraft. The No. 1 system is for the pilot and the No. 2system is for the copilot. The pilot and copilot canindependently select desired navigation aids. Duringoperation using navigation data from the dual radionavigation system (VOR, TACAN, ILS), the pilot’sflight director system receives information from the No.1 systems, and the copilot’s flight director systemreceives information from the No. 2 systems. The flightdirector systems present navigation data by means ofhorizontal situation indicators (HSI) and attitudedirector indicators (ADI) located on the pilot’s andcopilot’s instrument panels. The HSI display a pictorialview of the navigation situation as viewed from above.The ADI display a pictorial view of the navigationsituation as viewed from behind. The flight directorsystem has no on-off switch. Whenever power issupplied and the system circuit breakers are closed, thesystem is on. However, HSI and ADI data is not validuntil the INS and AHRS have completed their warm-upsequences. The INS requires approximately sevenminutes minimum reaction time.

The flight director system consists of the pilot’s andcopilot’s navigation selector panels, rate gyroscopetransmitter, power adequacy indicators, test switch,slaving switch, elevator position transmitters, flightdirector computer, HSI, and ADI .

Navigation Selector Panels

The navigation selectors (NS) provide the pilot andcopilot with switches for selecting various modes ofoperation of the flight director systems. One selectorpanel is located on each glare shield.

Rate Gyroscope Transmitter

The rate gyroscope transmitter provides turn rate signalsdirectly to the copilot’s ADI which are displayed on thecopilot’s ADI rate of turn indicator. The rate gyroscopetransmitter is located in the center underdeck avionicsrack.

Power Adequacy Indicators

The power adequacy indicators are located on the pilot’sand copilot’s instrument panel adjacent to the ADI. Theelectromechanical devices have a normally dark solidcolor face which displays a white RATE OFF (rate ofturn off) message when the turn rate signal to the relatedADI is unreliable.

Test Switch

A test switch is provided for each flight director system.The test switch for each system is located on the centerinstrument panel (Fig 2-7, 2 of 3).

Slaving Switch

If both pilots have selected an INS or radio mode, theslaving switch, located on the center console, enables thepilot to slave the copilot’s HSI to his HSI.

Elevator Position Transmitter

Each elevator position transmitter is located under theflight station floor at the base of a control wheel column.The elevator position transmitter provides a signaldetermined by the position of the control columns and istherefore related to the elevator position (Fig 2-7, 2 of3).

Flight Director Computer

The flight director computer receives and processesnavigational input signals for the purpose of providingsteering signals for display to the pilots . The pilot’sflight director computer (FDC) is located in the leftavionics compartment. The copilot’s FDC is located inthe center avionics equipment rack (Fig 2-7, 3 of 3).

Horizontal Situation Indicator (HSI)

The integrally lighted HSI are located on the pilot’s andcopilot’s instrument panels (Fig 2-7, 1 of 3).

Attitude Direction Indicator (ADI)

The integrally lighted ADI are located on the pilot’s andcopilot’s instrument panels (Fig 2-7, 1 of 3). Theypresent the view seen from aft of the aircraft lookingforward.


2-16

Navigation System

2-17

Fig 2-7, ( 1 of 3), Flight director system components


2-18

Fig 2-7, (2 of 3), Flight director system components

Navigation System

2-19

Fig 2-7, (3 of 3), Flight director system components

2-8 GLIDE SLOPE SYSTEM

The glide slope system is used as an aid to the pilot andcopilot for instrument landings. The glide slope systemconsist of two independent glide slope receivers thatshare a common antenna. Navigation system relatedcomponents such as the navigation selector panels,VHF/ COMM control panel, and attitude directorindicators (ADI) are also part of the glide slope system(Fig 2-8, 1 of 2).

Glide Slope Receiver

The glide slope receiver has been incorporated in theARN-147 VHF NAV reciever. Which is located in theleft avionics bay.

Glide Slope Antenna

The two glide slope systems share a common antenna.The antenna is located aft and below the search radarantenna in the nose radome (Fig 2-8 ,2 of 2).


2-20

Fig 2-8, ( 1 of 2), Glide slope system components

Navigation System

2-21

Fig 2-8, (2 of 2), Glide slope system components

2-9 MARKER BEACON SYSTEM (MKR BCN)

The marker beacon system informs the flight crew whenthe aircraft passes over a marker beacon ground station.The system provides the pilot and copilot with bothvisual and audio indications. Visual indications ofstation passage is provided by three indicator lights.Audio indication is by coded tones heard over theinterphone system in the flight crew’s headsets. Themarker beacon system depends upon ground basedstations. The interaction between ground stations and theaircraft system makes the marker beacon system work.

Two types of marker beacon ground stations exist: theairways type and the instrument landing system (ILS)type. The airways type is used as a navigation aid duringcross-country flights. The airways type transmits anaudio tone in Morse Code of its station identification,and also causes an indicator light to come on. Theinstrument landing system (ILS) type station providesthree marker reference points along the runwayapproach. Reference point identification is provided bythree different audio tones heard over the interphonesystem. A set of indicator lights will sequence on duringan approach.

The marker beacon system consists of a marker beaconreceiver, antenna, marker beacon indicator panels, and aMKR BCN HIGH/LOW switch (Fig 2-9) .

Marker Beacon Receiver

The marker beacon receiver is the same receiver as theARN-147 VHF NAV receiver.

Marker Beacon Antenna

The marker beacon system has one receiving antennawhich is located in the rear bottom area of the left mainlanding gear pod.

Marker Beacon Indicators

The pilot and copilot each have a marker beaconindicator. The indicators are located on the pilot’s andcopilot’s instrument panels (Fig 2-9).

MKR BCN HIGH/LOW Switch

The only control for the marker beacon system is thepilot’s marker beacon sensitivity switch. Placard MKRBCN HIGH/LOW, the switch is next to the pilot’smarker beacon indicator.


2-22

Navigation System

2-23

Fig 2-9, Marker beacon system components

2-10 IDENTIFICATION FRIEND OR FOE (IFF)

The IFF system provides automatic radar identificationof the aircraft when interrogated by surface or airborneradar sets. This APX-72 IFF is the airborne portion of theair traffic control (ATC) radar beacon system. The ATCradar beacon system is an international network thatcontrols and monitors movement of all aircraft. Stationsare strategically located to provide the most coverage inthe most densely traveled areas of the world. Also, thesystem enables friendly aircraft to identify themselvesapart from other friendly aircraft and provides a means oftransmitting a special coded signal known as anemergency reply. In addition to the identificationinformation, the reply signal can report the pressurealtitude of the aircraft.

The IFF system consists of an IFF control panel, aSCADC/IFF altitude data input switch panel, areceiver-transmitter (transponder), an IFF transpondertest set, two antennas, and an antenna switching unit. Thesystem also has a MODE 4 computer (Fig 2-10, 1 of 3thru 3 of 3).

IFF Control Panel

The IFF control panel is located on the center console,and contains the necessary controls to operate the IFFradar system (Fig 2-10, 1 of 3).

IFF Receiver-Transmitter

The IFF receiver- transmitter, located on the right side ofthe center avionics equipment rack, contains circuits thattransmit and receive radio frequency (RF) energy.

IFF Transponder Test Set

The IFF transponder test set is located in the centeravionics rack, above the receiver-transmitter.

IFF Antennas

The top IFF antenna is a blade type. It is located at FS404, buttock line 0. The bottom IFF antenna is flushmounted at FS 608 buttock line 0 (Fig 2-9, 3 of 3).

IFF Antenna Switching Unit

The IFF antenna switching unit is located in the centeravionics area and serves as a junction that has both thetop and bottom antenna (Fig 2-10, 2 of 3).

Mode 4 Computer

The Mode 4 computer is located on the left side of thecenter avionics rack (Fig 2-10, 3 of 3). The Mode 4computer, utilized only during Mode 4 operations,allows the receiver-transmitter to accept specialinterrogation signals and issue special reply signals.


2-24

Navigation System

2-25

Fig 2-10, (1 of 3), IFF system components


2-26


Navigation System

2-27


2-11 APS-133 COLOR WEATHER RADAR SYSTEM

The APS-133 color weather radar system has theprimary function of providing weather warninginformation. The weather radar system can also be usedfor long and short range terrain mapping, air-to-airmapping, and beacon interrogation. As a weatherwarning radar set, it provides a display of weatherinformation such as precipitation areas and stormclouds. As a terrain mapping radar, it provides a displayof prominent terrain features such as cities, shorelines,islands, and high ground as an aid to navigation. Thisinformation is presented as a color coded display on4-color multifunction displays mounted at the pilot’scenter instrument panel, and on SKE equipped aircraft,at the navigator’s position. The weather radar systemconsists of a receiver-transmitter, multifunction display,antenna, and radar control panel .


The radar system receiver-transmitter is located in theunderdeck center avionics rack (Fig 2-11, 2 of 2).

Multifunction Display

The multifunction display contains a 4-color cathoderay tube (CRT) (Fig 2-11, 1 of 2). The multifunctiondisplay indicator is located in the pilot’s centerinstrument panel. A second multifunction display islocated on the navigator’s table on aircraft equippedwith station keeping equipment (SKE).

Antenna

The weather radar antenna is mounted under the radomeon the pressure bulkhead at FS275 (Fig 2-12). Theantenna consists of an antenna pedestal and antennareflector which shape, tilt, rotate, and stabilize the radarbeam.

Radar Control Panel

The radar control panel is normally located in the pilot’scenter console. Alternate provisions for the radar controlpanel exist on the navigator’s instrument panel and areused on SKE equipped aircraft (Fig 2-11, 1 of 2).


2-28

Navigation System

2-29

Fig 2-11, (1 of 2), Radar system components


2-30

Fig 2-11, (2 of 2), Radar system components

Navigation System

2-31

Fig 2-12, Radar system antenna exterior controls

2-12 THE GROUND PROXIMITY WARNING SYSTEM (GPWS)

The ground proximity warning system (GPWS)provides the flight crew with visual and audio warning ofpossibly dangerous flight paths relative to the ground.GPWS issues visual warnings via lights on the pilot’sand copilot’s GPWS panels (Fig 2-13). The audiowarnings include synthesized voice commands heardover the GPWS loudspeaker and in the flight crew’sheadset if the interphone system is on. The warningmodes are active when the aircraft is flying between 50and 2450 feet above the ground. For this range ofaltitude, the GPWS monitor both the aircraft’s heightand change in height (altitude rate.) The system tries topredict when the aircraft is too close to the ground. When

a warning occurs, the flight crew must take correctiveaction (such as a pull-up). The warning stops only whenthe unsafe ground proximity situation no longer exists.Some warning modes can be inhibited for specialreasons (such as low-level flying). The GPWS monitorsits operational status as soon as initial power is applied.If the GPWS is inoperative, the GPWS INOPannunciator on the master caution panel will come on.The flight profile comparator (FPC) is located in thecenter avionics equipment rack under the flight deck.The FPC automatically monitors radar altitude,barometric altitude, and glide slope signals, as well aslanding gear and flap positions.


2-32

Fig 2-13, Ground proximity warning system

2-13 INERTIAL NAVIGATION SYSTEM (INS)

The INS is an all weather, world wide inertial navigationsystem with automatic radio position update capability.The system may be operated as an area navigationsystem or as a inertial navigation system (Fig 2-14). TheINS provides accurate global navigation withoutreference to ground signals. Navigation is accomplishedby the sensing and measuring of acceleration by theprecision gyro stabilized platforms in conjunction with adigital computer. Two of the platforms are used in thedual independent inertial navigation system (INS). Thethird platform provides an attitude and heading referencefor AHRS and a backup for INS. The INS employs acomputer and a highly sensitive, stable and precisemotion sensing mechanism to monitor direction,attitude, and velocity changes.

The INS provides a dual navigation system composed ofthe necessary sensors, computer processing andmonitoring displays to provide precision entirelyindependent of ground reference signals. Appropriatesystem outputs provide smooth, flyable commands tothe pilot, copilot, and autopilot. The fundamental abilityemployed in the system to accomplish the navigationfunction is the sensing and measurement of acceleration.

Inertial Navigation System Components

The INS primary system components consists of twonavigation units, one FSAS/INS CDU, (Fig 2-14, 3 of 4)two mode selector units, two battery units, (Fig 2-14, 2of 4) two BATAC units (which changes DC power to ACpower to operate the INS cooling fans),(Fig 2-14, 1 of 4),one pilot’s INS control unit, one navigator’s INS controlunit, (Fig 2-14, 4 of 4), two navigation selector panels(Fig 2-14, 2 of 4), and one INS CDU.

Both navigation units and BATAC are in the leftavionics equipment rack along with the INS junctionbox No. 2. Both battery units and selector relays are inthe center avionics equipment rack. The FSAS/INSCDU, the pilot’s INS control unit (PICU), and bothmode selector units (MSUs) are in the center console.The INS CDU, the navigator’s INS control unit (NICU),and INS junction box No. 1 are at the navigator’s station.The navigation selector panels are located on the glareshield above the center instrument panel.

Navigation System

2-33


2-34

Fig 2-14, (1 of 4), INS components

Navigation System

2-35



2-36


Navigation System

2-37


2-14 AUTOMATIC DIRECTION FINDER SYSTEM (ADF)

The ADF system is a radio receiver that has directiondetermining capability (Fig 2-15). The ADF systeminterfaces with the interphone/PA system, and thebearing distance heading indicators (BDHI) and relatedcontrols. The ADF provides the flight crew with a radioaid for information and navigation. The ADF is a radioreceiving system which can be tuned to stationsbroadcasting in the frequency range of 190 to 1750 KHz.The ADF can receive amplitude modulated (AM),continuous wave (CW) or unmodulated carrier signals.Many types of stations operate in this range, including

ground beacon stations, radio range stations, weatherstations, and standard AM radio stations. The ADF canalso be used as a radio compass with automatic andmanual direction finding. The relative bearing (angle)between aircraft heading and selected station isdisplayed at the pilot’s, copilot’s, and navigator’spositions on the BDHI. Each ADF systems consists of areceiver, control panel, sense antenna, antenna coupler,sense transmission line, loop antenna, quadrant errorcorrector unit, and loop transmission line (Fig 2-15, 1 of3 thur 3 of 3).


2-38

Fig 2-15, (1 of 3), ADF system components

Navigation System

2-39



2-40


2-15 VHF NAVIGATION SYSTEM

The VHF navigation system provides the aircraft withbearing information to selected ground stations. TheVHF navigation system provides the flight crew with aradio aid for VHF omni-range (VOR) and receiverlocalizer (LOC). VOR information is used as anavigational aid during cross-country flights. LOCinformation is used to provide lateral guidanceinformation as part of the instrument landing system(ILS).

The VHF navigation systems consist of two VHFnavigation receivers, two VHF COMM/NAV controlpanels, and they share a common set of special antennasand an antenna switch (Fig 2-16 and Fig 2-17).

VHF Navigation Receivers

The VHF navigation receivers are located in the centeravionics equipment rack under the flight deck (Fig 2-16,1 of 2).

VHF COMM/NAV Control Panels

The pilot and copilot each have a VHF COMM/NAVcontrol panel located on the center console. The controlpanels are used for mode selection and tuning. Anumeric display indicates the selected channelfrequency (Fig 2-16, 2 of 2).

VHF NAV VOR Tail Antennas

The VHF NAV VOR tail antennas are located on eachside of the vertical stabilizer (Fig, 2-16, 1 of 2).

VHF NAV LOC Nose Antenna

The VHF NAV LOC nose antenna is located in the noseradome (Fig 2-16, 1 of 2).

VHF NAV Antenna Switch

The VHF NAV antenna switch is located in the leftunderdeck avionics rack.

Navigation System

2-41


2-42

Fig 2-16, (1 of 2), VHF navigation system components

Navigation System

2-43

Fig 2-16, (2 of 2), VHF navigation system components


2-44

Fig 2-17, VHF navigation system related components

2-16 AN/ARN-118 TACTICAL AIRCRAFT NAVIGATION SYSTEM

(TACAN SYSTEM)

The AN/ARN-118 airborne TACAN system consists ofa TACAN receiver-transmitter, TACAN control, adigital-to-analog adapter, and a shock mount. Thesystem provides the pilot and flight control systems withaccurate and reliable navigational information on bothX and Y TACAN channels. A self-contained, automaticantenna switch allows the system to operate with eithersingle or dual TACAN antennas (Fig 2-18, 2 of 2).

Since the TACAN system operating limit is line of sight,the actual operating range is dependent on aircraftaltitude. The system is designed to prevent 40-degreeerror lock-on, co-channel interference, and lock-on tofalse or incorrect signals. The system uses advanceddigital circuitry to reduce space and weight. TheAN/ARN-118 airborne TACAN system produces

slant-range distance, relative bearing, course deviation,to-from, and audio identification information for use bythe flight crew. The distance, bearing, and audioidentification information indicates the location of acomplementary surface station with respect to theaircraft. The system produces distance information withrespect to a similarly equipped aircraft or distance,bearing, and course deviation information with respectto a suitably equipped, cooperating aircraft as anauxiliary function.

The receiver-transmitter (RT) unit and instrumentationcoupler are in the left avionics equipment rack. Thecontrol panels are installed on the center console (Fig2-18, 1 of 2). The bottom antenna for each system isflush mounted on the bottom center line of the fuselage.

Navigation System

2-45

Fig 2-18, (1 of 2), TACAN system components


2-46

Fig 2-18, (2 of 2) TACAN system components

2-17 AN/APN 169C INTRAFORMATION POSITIONING SYSTEM

(SKE)

The AN/APN 169C intraformation positioning system,or station keeping equipment (SKE), is an airborneelectronic system providing the flight crew withpositioning information (Fig 2-19). The system providesthe capability for close formation flying with otheraircraft also equipped with SKE. The SKE is aspecialized light weight system which allows up to 36aircraft to maintain fixed separation between oneanother in formation. The aircraft can locate each otherand identify the leader aircraft during day or night flightsunder all weather conditions. SKE performs four basicfunctions. First, the system displays station keepinginformation (relative range and azimuth positioninginformation of other cooperating SKE equipped aircraftin the formation) on the pilot’s multifunction display.Second, the system indicates track-while-scan (TWS)positioning data (in-track, cross-track, and altitude)relative to the selected leader aircraft on the pilot’s andcopilot’s ADI and the pilot’s and navigator’s relativerange indicators. Third, the system has an integral silent(no voice) signaling capability for the transfer of discretedata. Some of the data are predefined commands such asturn left, fly up, or increase airspeed. Numerical data,such as desired airspeed or delay time, can be sent to adigital readout indicator. These signals coordinatechanges to the flight path of the formation. Last, SKE hasan audio and visual alarm warning system to signal thepresence of SKE equipped aircraft intruding within aselectable zone. The SKE can be operated in conjunctionwith a ground based zone marker radar set for airdelivery missions. The zone marker is positioned(usually air dropped) in the drop zone area. The zonemarker is deployed as operating ground equipment.

The following components are part of the intraformationpositioning set (Fig 2-19, 1of 4 thru 4 of 4): primarycontrol, secondary control, flight command indicator,flight command indicator repeater, relative rangeindicator, audio amplifier, auxiliary interface unit, radarreceiver-transmitter, coder-decoder, electricalequipment rack, two antennas, antenna pedestal, andassociated components.

Primary Control

The primary control is located on the navigator’sinstrument panel (Fig 2-19, 1 of 4). The primary controlcontains switches, indicators, and the BITE (built in testequipment) controls for the SKE system.

Secondary Control

The secondary control is located on the navigator’sinstrument panel (Fig 2-19, 1 of 4). The secondarycontrol contains controls and switches to select thetrack-while-scan (TWS) and proximity warningparameters for SKE operation.

Flight Command Indicator (FCI)

The flight command indicator is located on thenavigator’s instrument panel. The FCI contains the frontpanel controls and indicators that initiate thetransmission and reception of three-decimal digits ofdata, such as true airspeed.

Flight Command Repeater (FCR)

The FCR is located above the pilot’s center instrumentpanel (Fig 2-19, 1 of 4). The pilot’s FCR containsindicators that display the information being sent fromor received by the FCI.

Relative Range Indicator (RRI)

There is two RRIs, one located on the pilot’s centerinstrument panel and the other on the navigator’sinstrument panel (Fig 2-19, 1 of 4). The RRI provide avisual presentation of own aircraft in-track location withrespect to the selected range to the leader aircraft.

Audio Amplifier

The audio amplifier contains an audio frequency poweramplifier to drive the aircraft audio alarm speaker. Theaudio amplifier is utilized for audio proximity warningindications of the SKE, and to signal upon reception offlight commands on the FCI. The SKE audio amplifier islocated in the center avionics equipment rack (Fig 2-19,3 of 4).

Auxiliary Interface Unit (AIU)

The AIU is a signal data converter. The AIU and it’smount are located in the underdeck avionics equipmentrack (2-19, 2 of 4).

Navigation System

2-47


2-48

Fig 2-19, (1 of 4), SKE system components

Navigation System

2-49



2-50


Navigation System

2-51



The receiver-transmitter is located next to thecoder-decoder in an overhead equipment rack in thecargo area (Fig 2-19, 2 of 4).

Coder-Decoder

The coder-decoder is located next to thereceiver-transmitter in an overhead equipment rack inthe cargo area (Fig 2-19, 2 of 4).

Electrical Equipment Rack

The equipment rack, located overhead in the cargo area,is a combination junction box and mounting frame.

Antenna

The directional antenna is center mounted on a pedestalassembly under the aircraft near the main landing gear(Fig 2-19, 3 of 4).

Antenna Pedestal

The antenna pedestal is what the antenna is mounted toand which allows it to move 360 degrees.


2-52

2-18 FUEL SAVINGS ADVISORY SYSTEM (FSAS)

The fuel savings advisory system (FSAS) is anequipment change to the aircraft to provide an integratedgroup of line replacement units (LRUs). FSAS enablesthe flight crew to achieve military flight profiles withminimum fuel consumption. FSAS will also provideauxiliary displays of target engine pressure ratio (EPR),altitude (ALT), and airspeed (IAS), way point data,ground speed, and true heading to the display interfaceunit (DIU) to be displayed on the multifunction display(MFD). FSAS also provides both audio and visualwindshear/altitude warnings. The fundamental abilityemployed in FSAS is the computation of an optimal

flight path based on two factors: amount of fuel savedand amount of time saved. A flight index number of 0 to200 is selected with 0 representing most fuel consciousand 200 representing most time conscious.

The following are the primary components of the fuelsavings advisory system (Fig 2-20, 1 of 3 thur 3 of 3):fuel savings computer (FSC), fuel savings advisorysystem/inertial navigation system control display unit(FSAS/INS CDU), display interface control unit(DICU), and display interface unit (DIU).

Navigation System

2-53

Fig 2-20, (1 of 3), FSAS components


2-54

Fig 2-20, (2 of 3), FSAS components

Navigation System

2-55

Fig 2-20,(3 of 3), FSAS components


2-56

SECTION III INDICATING AND

RECORDING SYSTEM

This section consists of information extracted from T.O. 1C-141B-2-31GS-00-1.


1. The purpose of the indicating and recording system

2. The function of

• cockpit voice recorder

• digital flight data recorder

• takeoff warning


• cockpit voice recorder

• digital flight data recorder

Indicating and Recording System

3-1


3-2

3-1 INDICATING AND RECORDING SYSTEM

The indicating and recording system of the C-141Baircraft includes the cockpit voice recorder (CVR),digital flight data recorder (DFDR), and takeoff warningsystem. The recorders provide data about the aircraft and

its systems for accident investigation data, stress data,and aircraft usage data. The takeoff warning systemprovides a visual cue that the aircraft is configured fortakeoff.

3-2 COCKPIT VOICE RECORDER (CVR) SYSTEM

The CVR system uses a magnetic tape recorder toautomatically record the latest 30 minutes ofcommunications between flight station crew members.The recorder is located in the aft upper deck and iscolored international orange. Inputs through theinterphone system, either received or transmitted, by thepilot, copilot, or flight engineer are recorded. A CVRcontrol unit in the overhead trim, aft of the pilot’soverhead panel, contains an area microphone. Themicrophone picks up any audio in the flight station,amplifies the signal, then routes the result to the recorder.A test switch on the control unit initiates the test circuit inthe recorder. Sequential tones of the four recordingchannels can be monitored by a headset plugged into thecontrol unit. A meter also provides a visual indication ofthe test. The CVR system uses an external powerinterlock to prevent previously recorded data from beingerased while the aircraft is on the ground and externalpower is connected. A CVR TEST switch can be used tobypass the interlock for maintenance purposes. TheCVR system operates automatically without affectingany of the aircraft systems.

ComponentsThe CVR system consists of the cockpit voice recorder(CVR), CVR control unit, and a CVR TEST switch.

Cockpit Voice Recorder (CVR)The CVR is located in the aft upper deck and is coloredinternational orange. The insulated case protects therecorded data from shock, fire, and salt water that mayresult from an aircraft accident (Fig 3-2).

Cockpit Voice Recorder (CVR) ControlUnitThe CVR control unit provides remote audio monitoringand test of the CVR. The control unit also contains thearea microphone for recording flight station audio. Thecontrol unit is mounted in the trim panel aft of the pilot’soverhead panel (Fig 3-2).

Cockpit Voice Recorder TEST SwitchThe TEST switch is provided to ensure previouslyrecorded flight data is not overwritten on the ground. Anexternal power lock is provided. The TEST switch islocated on the center avionics rack in the right avionicscompartment (Fig 3-1).


3-3

Fig 3-1, Cockpit voice recorder test switch


3-4

Fig 3-2, Cockpit voice recorder components


3-5

Fig 3-3, Digital flight data recorder component location

3-3 DIGITAL FLIGHT DATA RECORDER (DFDR) SYSTEM

The digital flight data recorder (DFDR) system recordsthe latest 25 hours of aircraft flight performance data.Seventeen flight performance parameters areautomatically recorded. The parameters are:acceleration (vertical, lateral, and longitudinal), rudderposition, altitude, airspeed, heading, pitch attitude, rollattitude, pitch trim, engine pressure ratio (EPR), flapposition, spoiler position, thrust reversers (extended andlocked), landing gear position, radio keying, and aircraftidentification. Operation of the DFDR system is

automatic and has no affect on flight operation of theaircraft. The only indication available to the flight crewfor system status is a FLT REC INOP annunciator light.An external power interlock prevents data from beingoverwritten while the aircraft is on the ground.

The DFDR is located in the aft upper deck next to thecockpit voice recorder (CVR). The DFDR is coloredinternational orange. The insulated case protects therecorded data from shock, fire, and salt water that mayresult from an aircraft accident (Fig 3-3).

3-4 TAKEOFF WARNING SYSTEM

The takeoff warning system provides a visual indicationthat certain aircraft systems are in their normal positionprior to takeoff. A green TAKEOFF warning light comeson when the systems are normal and the button on thehydraulic pitch trim lever is depressed. When thesystems are in their normal position, the followingconditions are met: isolated AC avionics bus is powered,

isolated AC bus is powered, main DC buses No. 1 andNo. 2 are powered, spoilers are closed and locked, allthrust reversers are stowed and locked, flaps are intakeoff position, autopilot is off, all external doors areclosed and locked, and the spoiler handle is armed.

3-5 LIFE HISTORY RECORDER SYSTEM (LHRS)

This system has been deactivated.


3-6

SECTION IV AUTOMATIC FLIGHT

CONTROL SYSTEM



1. The purpose of the automatic flight control system

2. The function of

• autopilot system

• yaw damper system

• automatic throttle system

• fuel saving advisory system

• all weather landing system


• autopilot system

• yaw damper system

• automatic throttle system

• fuel saving advisory system

• all weather landing system

Automatic Flight Control System

4-1


4-2

4-1 AUTOMATIC FLIGHT CONTROL SYSTEM

The automatic flight control system (AFCS) of theC-141B aircraft includes the autopilot (A/P) system,yaw damper (Y/D), automatic throttle system (ATS),fuel saving advisory system (FSAS), and the all weather

landing system (AWLS) This system provides a meansfor automatically controlling the aircraft in normalflight. It also guides the aircraft to runways for landings.

4-2 AUTOPILOT

The autopilot provides the means by which the aircraft iscontrolled automatically in flight. The autopilot canguide the aircraft to the runway for landings. The actionof the autopilot is smooth since autopilot signals forcorrective action are directly proportional to the amountof displacement. Coordinated turns are made in allmodes of operation.

Autopilot Controls and Indicators

The automatic flight control system (AFCS) panel islocated on the pilot’s center console. With this panel, thepilot can control the desired mode engagement. Inaddition, he can change altitude or make coordinatedturns without disengaging the autopilot.

AFCS Control Panel Procedures

The autopilot switch on this panel is used to engage theautopilot. It can also be used to disengage the autopilot.If the autopilot disconnect switch on either control wheelis depressed, the autopilot will disengage and theautopilot switch will return to the off position. The yawdamper system does not disengage when the autopilot isdisengaged.

Control Wheel Steering

Control wheel steering (CWS) is incorporated into thepilot’s control wheel only. It allows aircraft attitudecontrol without disengaging the autopilot. Control wheelsteering mode is selected by placing the CWS selectorswitch, on the AFCS control panel, forward to the CWSposition. A wheel pressure of more than 2.5 pounds willactivate the system.

Autopilot Mode Selection

The various autopilot modes have definite priorities andcompatibility. The term compatibility is used to indicatethat two different modes can be selected at the sametime. For example, altitude hold and roll CWS can be

selected at the same time. Some modes have priorityover other modes. If a higher priority mode is selected,the lower priority mode will drop out of the controlcircuit. The interlock system prevents the pilot fromselecting a low priority mode when a higher prioritymode has already been selected. Below are the differentmodes of operation:

Autopilot Operational Modes

The autopilot modes are selected by switches on theAFCS control panel (Fig 4-1). Selection of navigationaids used with the autopilot are made on the copilot’snavigation selector panel.

Heading Hold Mode

This is a primary mode. In this mode, the autopilot usesthe existing aircraft heading at time of switchengagement as the heading reference. This mode isselected when the AUTOPILOT switch (AFCS panel) isturned ON. It is not in operation when CWS or theTURN controller is being used.

Altitude Hold Mode

This mode holds the aircraft at the reference altitudeexisting at time of engagement.

Mach Hold Mode

This mode maintains the aircraft at the Mach existing atthe time of engagement. The Heading Hold Mode isretained. Other heading modes may be selected.

Pitch Off Mode

Selecting PITCH OFF with the ALT HOLD/PITCHOFF switch turns off pitch control only. The rest of theautopilot is still engaged. The pilot and/or copilot has tofly the pitch axis manually in this condition.


4-3

Lateral Off Mode

Selecting LAT OFF with the NAV SEL/LAT OFFswitch turns off lateral steering only. The rest of theautopilot is still engaged. The pilot and/or copilot fliesthe lateral axis in this mode.

Navigation Select Operation

Basic autopilot functions may be supplemented bysignals from the navigation systems installed on theC-141B. Selection of the desired aid is made on thecopilot’s navigation selector panel.

Heading Select Mode

In this mode, the autopilot maintains the aircraft on theheading set with the heading marker on the copilot’sheading set indicator (HSI).

(VOR)(ILS) Mode

This mode is placed in service by depressing theVOR/ILS button on the copilot’s navigation selector

panel, and placing the (HDG SEL/NAV), button toNAV. After these two have been set, set the (NAVSEL/LAT OFF) switch (AFCS panel) to NAV SEL.When intercept heading is established, the autopilot willremain responsive to the heading marker until the VORcourse deviation is one dot deflection, then thenavigation aid takes over control.

INS Select Mode

To use INS, select INS 1 or INS 2 on the copilot’snavigation selector panel. On the AFCS panel, place theNAV SEL/LAT OFF switch to NAV SEL. The autopilotwill then intercept and track the active INS course if thecopilot’s HDG SEL/NAV button is in NAV.

Fuel Savings Advisory System Mode(FSAS)

To use the FSAS, select the mode of operation in theFSAS, press the lighted engage button on the FSASCDU, and then take the G/S/VER NAV switch to theVER NAV position. The aircraft will now maintain orfly the mode selected in the FSAS.


4-4


4-5

Fig 4-1, Autopilot components

4-3 YAW DAMPER SYSTEM

The Automatic Flight Control System (AFCS) includesa yaw damper system. This yaw damper system operatesindependently of the autopilot portion of the AFCS andis electrically isolated from the autopilot. The yawdamper system controls the rudder in all modes ofautomatic operation.

Theory of Operation

Yaw damper engagement means that the rudder controlsurface is being positioned by the rudder servos of theAFCS. This engagement is accomplished by placing theyaw damper switch on the yaw damper panel to ON. Therudder servos have automatically assumed the positionof the aircraft prior to this, resulting in a smooth takeoverby the yaw damper. If the autopilot is not operating, yawdamper rate gyros No. 1, No. 2, and No. 3 sense anydeviation of the aircraft in the yaw axis.

The yaw damper control panel is located on the pilot’scenter console. To turn on the yaw damper, place theOFF-ON switch to ON.

Warning System

If one of the yaw rate gyros or servos malfunctions, themonitor system will turn on the YAW DAMPERFAULT light located on the pilot’s annunciator panel. Ifmore than one of the yaw rate gyros and/or servosmalfunction, it is detected by the monitor circuit whichturns on the YAW DAMPER INOPERATIVE lightslocated on the pilot’s and copilot’s instrument panelsand automatically disengages the YAW DAMPER.

Testing

The yaw damper test circuit is wired through thetouchdown relays, so that it is rendered inoperativewhenever the aircraft is in flight. Testing isaccomplished from the yaw damper control panellocated on the pilot’s center console.

4-4 AUTOMATIC THROTTLE SYSTEM (ATS)

The ATS is used to maintain indicated airspeed at aconstant preselected value. It also provides an automaticthrottle retard rate when the system is used in an AWLSapproach. The ATS engaged switch becomes the aerialrefueling boom disconnected switch when the AERIALREFUEL MASTER switch is positioned to ON.

The components and warning lights of the automaticthrottle system consist of an airspeed trim assembly,ATS computer, motor/generator assembly, clutch packassembly, system arm and engaged switches, an AUTOTHROTTLE friction light located below the AUTOTHROTTLE switch, and a THROTTLE light on theAWLS fault identification panel (Fig 4-2, 1 of 2).

Airspeed Trim Assembly

The airspeed trim assembly consist of a positiontransducer which is controlled by a spring-loaded,detented, auto throttle trim thumb wheel. It is located onthe center console to the right of the throttle frictionknob. The airspeed trim assembly permits changes in theindicated airspeed, after the ATS mode has beenengaged, without first disengaging.

ATS Computer/Amplifier Unit

The ATS computer/amplifier unit is located in theavionics underdeck rack. The computer receives inputsfrom both central air data computers, the testprogrammer logic computer, the airspeed trimassembly, the flare computer, and feedback signals fromthe position sychros in the motor/generator which drivesthe clutch pack.

Motor Generator Assembly

The motor/generator assembly is mounted on the clutchpack (Fig 4-2, 2 of 2). The components consist of aservomotor and a rate generator. Also included are aservo position synchro transmitter and a 10:1 reductiongear train.

Clutch Pack Assembly

The clutch pack is in the avionics underdeck rack (Fig4-2, 2 of 2). The clutch pack consist of a solenoidoperated friction disk clutch gear train, and a positionsychro transmitter. Also included are maximum andminimum throttle position switches, idle disconnectswitches, and four throttle cable quadrants.


4-6


4-7

Fig 4-2, (1 of 2), Auto throttles components


4-8

Fig 4-2, (2 of 2), Auto throttles components

4-5 FUEL SAVINGS ADVISORY SYSTEM (FSAS)

The FSAS is coupled to the automatic throttle system toprovide aircraft speed and engine pressure ratio (EPR)control during climb and cruise phases of flight. In theauto throttle system, the existing input signals from thevelocity error, speed trim, pitch attitude, and throttleposition will not be utilized when the ATS is coupled tothe FSAS. The only remaining signal into the ATS servoamplifier will be from the servo motor generator. TheFSAS throttle rate command will use the same wires asthe throttle feedback. Therefore, the throttle inner loopwill be a throttle rate control system.

FSAS INS Navigation Aiding

The FSAS system makes inertial navigation less of awork load requirement on the aircrew by increasing thenumber of waypoints that can be programmed beforeand after takeoff. The FSAS has a total memorycapability of 40 waypoints.

FSAS Told Card Information

The FSAS completes a takeoff and landing data (TOLD)card for the aircrew when the conditions affecting thetakeoff and type of takeoff desired are selected andinserted into the system. The TOLD card can be based ona reduced EPR or total rated thrust (TRT) takeoff with orwithout bleed air penalties. When performanceparameters are exceeded, the system will warn theaircrew to that effect. For example, “Gross weight islimited by critical field length.” If the takeoff is stillrequired with the conditions present, the system willcompute a new take-off TRT based on setting TRT priorto brake release.

Flight Planning

The FSAS system helps the pilots compute the exact fuelrequired to fly either an entire mission or a leg segment.The winds and temperatures between waypoints, desiredflight altitude, climb, cruise and descent profiles to beflown can be entered to determine the fuel requirementsfor an entire flight. The system will also display fuel,time, and range remaining based upon presentconditions or climbing to an optimum altitudes.

Flight Profiles

Once the aircraft is 1,500 feet above ground level, thefull capabilities of the FSAS can be taken advantage of.

If the autopilot and auto throttle systems are turned onand engaged with the FSAS, the aircraft canautomatically climb in one of four modes up to a selectedaltitude.

FSAS INS Radar Displays

The FSAS system will allow for the display ofinformation on the AN/APS-133 radar scope. Some ofthe information that can be displayed, is as follows:reference ground speed, actual ground speed, optimumaltitude, airspeed or mach, INS true heading, enginepressure ratio, and variable way-point latitude andlongitude. As an aid to the aircrew, the FSAS will alsodisplay on the radar scope a scale display of the aircraft’sposition with course lines and way point TACANsymbols. This is to show the aircraft’s present positionand the present position of relative points of interest.

FSAS Components

The four major components of the FSAS are the FSAScomputer, the FSAS display interface unit, the displayinterface control unit, and the FSAS/INS control displayunit.

FSAS Computer

The FSAS computer, on the upper right side of theforward equipment rack in the left hand underdeck area,handles all calculations and displays of the FSAS.

FSAS Display Interface Unit

The FSAS display interface unit (DIU), on the left sideof the center avionics rack in the left hand underdeckarea, converts and then displays information from theFSAS computer to the AN/APS-133 radar scope.

FSAS Display Interface Control Unit

The display interface control unit (DICU), in the centerrow of the pilot’s center console just aft of the autopilotcontrol panel, controls the FSAS system’s operation, onor off, and allows the selection of the different displaysand information to be presented on the AN/APS-133radar scope.


4-9

FSAS Control Display Unit

The FSAS/INS control display unit (CDU), controls anddisplays all information provided by the FSAS to the

aircrew. This CDU is also the pilot’s only direct linkwith the INS system.

4-6 ALL WEATHER LANDING SYSTEM (AWLS)

The AWLS is used for landings during any weathercondition, good or bad. The AWLS includes built in testequipment (BITE) to perform enroute and prelandself-testing of all critical component functions and thefunctions of associated systems while in flight and priorto landing. The AWLS provides either automatic controlor guidance for manual control of the aircraft duringnormal landing approaches down to a minimum decisionaltitude (MDA) of 100 feet above the runway. If therunway is in sight at MDA, the AWLS continues toprovide either automatic control or guidance for manualcontrol through the flare and throttle retard maneuvers tolanding. At approximately 15 feet, the pilot manuallyaligns the aircraft with the runway (decrabs). If therunway is not visible at the MDA, a missed landingmaneuver can be initiated by pressing the GO AROUNDSWITCH on either control wheel and advancing thethrottles. In the ROTATION/GO AROUND (R/GA)mode, the R/GA computer provides pitch steeringguidance to the attitude director indicator as the pilotsmanually control the aircraft. The autopilot providesautomatic control of the flight control surfaces duringautomatic AWLS landings. The flight director systemsupplies visual guidance to the pilots as they manuallycontrol the flight control surfaces for manual AWLSlandings. The AWLS master caution system indicatesfaulty channels or modes of the AWLS that are detectedduring inflight self-test programs and also confirmsproper performance of AWLS functions at critical stagesof AWLS landing operations (Fig 4-3).

AWLS Components

The AWLS components include the test programmerand logic computer (TPLC), (Fig 4-4, 2 of 4), flarecomputer, rotation/go around (R/GA) computer,autopilot coupler, AWLS master caution system, four

vertical accelerometers, one horizontal accelerometer,(Fig 4-4, 1 of 4) an arming switch, a test switch, and aAWLS junction box (Fig 4-4, 2 of 4). The AWLS mastercaution system includes two flight progress andauto-man display panel, and a master caution controlbox. A self-testing capability is incorporated into theAWLS components and the system (Fig 4-4, 4 of 4). Thisbuilt-in-testequipment (BITE) is used to test the systemand components while in flight and also during groundcheckout procedures.

AWLS Test Programmer and LogicComputer (TPLC)

The TPLC is located in the underdeck electronicequipment area. The TPLC provides a warningindication during enroute and preland test during AWLSoperation.

Rotation-Go-Around Computer (RGA)

The RGA computer is located in the underdeck avionicsequipment area. The RGA computer is used to providepitch steering commands to both ADIs during takeoff orgo around. NOTE: Due to reliability and maintainabilitythe RGA system has been deactivated. There is a jumperplug install to allow operation of the copilots VVI (IAWHQ AMC MSG 0112052 Dec 93).

Flare Computer

The flare computer is located in the underdeckequipment area. The computer provides output signalsto the autopilot and FDS for executing an automatic ormanual flare.


4-10


4-11

Fig 4-3, All weather landing system control lights


4-12

Fig 4-4, (1 of 4), All weather landing system components


4-13



4-14



4-15



4-16

Documents

C-141B COMMAND AIRCRAFT SYSTEMS TRAINING (CAST) …