TECHNICAL MANUAL WARNING DATA …TM 55-1510-219-10 TECHNICAL MANUAL WARNING DATA OPERATOR'S MANUAL TABLE OF CONTENTS FOR ARMY RC-12D AIRCRAFT INTRODUCTION DESCRIPTION AND This copy

TM 55-1510-219-10

TECHNICAL MANUAL WARNING DATA

OPERATOR'S MANUAL TABLE OF CONTENTSFOR

ARMY RC-12D AIRCRAFT INTRODUCTION

DESCRIPTION ANDThis copy is a reprint which includes current OPERATONpages from Change 1.

AVIONICS

MISSION EQUIPMENT

OPERATING LIMITS ANDRESTRICTIONS

WEIGHT/BALANCE ANDLOADING

PERFORMANCE DATA

NORMAL PROCEDURES

EMERGENCY PROCEDURES

REFERENCES

ABBREVIATIONS AND TERMS

ALPHABETICAL INDEX"Approved for public release; distribution is unlimited."

*This TM 55-1510-219-10 supersedes TM 55-1510-219-10dated 25 May 1985 including all changes.

HEADQUARTERS, DEPARTMENTOF THE ARMY31 May 1991

URGENTTM 55-1510-219-10

C3

CHANGE

NO. 3

HEADQUARTERSDEPARTMENT OF THE ARMY

WASHINGTON, D.C., 12 May 1998

OPERATOR’S MANUALFOR

ARMY RC-12D AIRCRAFT

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited

TM 55-1510-219-10, 31 May 1991, is changed as follows:

1. Remove and insert pages as indicated below. New or changed text material is indicated by verticalbar in the margin. An illustration change is indicated by a miniature pointing hand.

Remove pages Insert pages

i and ii58.1/(5-8.2 blank)5-9 and 5-108-25 and 8-268-26.1/(8-26.2 blank)-------9-3 through 9-10-------INDEX-3 and INDEX-4

2. Retain this sheet in front of manual for reference purposes.

By Order of the Secretary of the Army:

Administrative Assistant to theSecretary of the Army

i and ii5-8.1 and 5-8.25-9 and 5-108-25 and 8-268-26.1 and 8-26.2(9-2.1 blank)/9-2.29-3 through 9-109-10.1/(9-10.2 blank)INDEX-3 and INDEX-4

DENNIS J. REIMERGeneral, United States Army

Chief of Staff

DISTRIBUTION:To be distributed in accordance with Initial Distribution Number (IDN) 310121, requirements for TM 55-

1510-219-10.

URGENT

TM 55-1510-219-10C2

CHANGE HEADQUARTERSDEPARTMENT OF THE ARMY

NO. 2 WASHINGTON, D.C., 12 December 1994

Operator's Manual

For


DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.


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GORDON R. SULLIVANGeneral, United States Army

Chief of Staff

MILTON H. HAMILTONAdministrative Assistant to the

Secretary of the Army07780

DISTRIBUTION:To be distributed in accordance with DA Form 12-31-E, block no. 0121, requirements for TM 55-1510-219-10.

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TM 55-1510-219-10C 1

CHANGE HEADQUARTERSDEPARTMENT OF THE ARMY

NO. 1 WASHINGTON, D.C., 7 August 1992

Operator's ManualFor



1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical barin the margin. An illustration change is indicated by a miniature pointing hand.

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- - - - 5-8.1/5-8.25-9 and 5-10 5-9 and 5-108-15 through 8-18 8-15 through 8-188-25 and 8-26 8-25 and 8-26- - - - 8-26.1/8-26.2



GORDON R. SULLIVANGeneral, United States Army

Official: Chief of Staff

MILTON H. HAMILTONAdministrative Assistant to the

Secretary of the Army

DISTRIBUTION:To be distributed in accordance with DA Form 12-31-E, block no. 0121, -10 & CL maintenance requirements for TM

55-1510-219-10.

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.

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TM 55-1510-219-10

WARNING PAGE

Personnel performing operations, procedures and practices which are included or implied in this technical manualshall observe the following warnings. Disregard of these warnings and precautionary information can cause seriousinjury or loss of life.

NOISE LEVELS

Sound pressure levels in this aircraft during some operating conditions exceed the Surgeon General's hearingconservation criteria, as defined in TM MED 501. Hearing protection devices, such as the aviator helmet or ear plugsshall be worn by all personnel in and around the aircraft during its operation.

STARTING ENGINES

Operating procedures or practices defined in this Technical Manual must be followed correctly. Failure to do somay result in personal injury or loss of life.

Exposure to exhaust gases shall be avoided since exhaust gases are an irritant to eyes, skin and respiratorysystem.

HIGH VOLTAGE

High voltage is a possible hazard around AC inverters, ignition exciter units, and strobe beacons.

USE OF FIRE EXTINGUISHERS IN CONFINED AREAS

Monobromotrifluoromethane (CF3Br) is very volatile, but is not easily detected by its odor. Although non toxic, it must beconsidered to be about the same as other freons and carbon dioxide, causing danger to personnel primarily by reductionof oxygen available for proper breathing. During operation of the fire extinguisher, ventilate personnel areas with freshair. The liquid shall not be allowed to come into contact with the skin, as it may cause frostbite or low temperature burnsbecause of its very low boiling point.

VERTIGO

The strobe/beacon lights should be turned off during flight through clouds to prevent sensations of vertigo, as a result ofreflections of the light on the clouds.

CARBON MONOXIDE

When smoke, suspected carbon monoxide fumes, or symptoms of lack of oxygen (hypoxia) exist, all personnel shallimmediately don oxygen masks, and activate the oxygen system.

FUEL AND OIL HANDLING

Turbine fuels and lubricating oils contain additives which are poisonous and readily absorbed through the skin. Do notallow them to remain on skin.

SERVICING AIRCRAFT

When conditions permit, the aircraft shall be positioned so that the wind will carry fuel vapors away from all possiblesources of ignition. The fueling unit shall maintain a distance of 20 feet between unit and filler point. A minimum of 10feet shall be maintained between fueling unit and aircraft.

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Prior to refueling, the hose nozzle static ground wire shall be attached to the grounding lugs that are locatedadjacent to filler openings.

SERVICING BATTERY

Improper service of the nickel-cadmium battery is dangerous and may result in both bodily injury and equipmentdamage. The battery shall be serviced in accordance with applicable manuals by qualified personnel only.

Wear rubber gloves, apron, and face shield when handling batteries. If corrosive battery electrolyte (potassiumhydroxide) is spilled on clothing or other material, wash immediately with clean water. If spilled on personnel,immediately start flushing the affected area with clean water. Continue washing until medical assistance arrives.

JET BLAST

Occasionally, during starting, excess fuel accumulation in the combustion chamber causes flames to be blownfrom the exhausts. This area shall be clear of personnel and flammable materials.

RADIOACTIVE MATERIAL

Instruments contained in this aircraft may contain radioactive material (TB 55-1500-314-25). These items presentno radiation hazard to personnel unless seal has been broken due to aging or has accidentally been broken. If seal issuspected to have been broken, notify Radioactive Protective Officer.

RF BURNS

Do not stand near the antennas when they are transmitting.

OPERATION OF AIRCRAFT ON GROUND

At all times during a towing operation, be sure there is a man in the cockpit to operate the brakes.

Personnel should take every precaution against slipping or falling. Make sure guard rails are installed when usingmaintenance stands.

Engines shall be started and operated only by authorized personnel. Reference AR 95-1.

Insure that landing gear control handle is in the DN position.

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TM 55-1510-219-10

TECHNICAL MANUAL HEADQUARTERSDEPARTMENT OF THE ARMY

NO. 55-1510-219-10 WASHINGTON, D.C., 31 May 1991

Operator’s Manual

ARMY MODEL RC-12D AIRCRAFT

REPORTING OR ERRORS AND RECOMMENDING IMPROVEMENTS

You can help improve this manual. If you find any mistakes or if you know of any way to improve the procedures,please let use know. Mail your letter, DA Form 2028 (Recommended Changes to Publications and Blank Forms), orDA Form 2028-2 located in the back of this manual directly to: Commander, U.S. Army Aviation and MissileCommand, ATTN: AMSAM-MMC-LS-LP, Redstone Arsenal, AL 35898-5230. A reply will be furnished directlyto you. You may also send in your comments electronically to our E-mail address at <[email protected]>, orby fax at (205) 842-6546 or DSN 788-6546. Instructions for sending an Electronic DA Form 2028 may be found atthe back of this manual immediately preceding the hard copy DA Forms 2028.

TABLE OF CONTENTS

CHAPTER 1. INTRODUCTIONCHAPTER 2. AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION

Section I. AircraftII. Emergency equipment

III. Engine and related systemsIV. Fuel systems

V. Flight controlsVI. Propellers

VII. Utility systemsVIII. Heating, ventilation, cooling, and environmental control system

IX. Electrical power supply and distribution systemX. Lighting

XI. Flight instrumentsXII. Servicing, parking, and mooring

CHAPTER 3. AVIONICSSection I. General

II. CommunicationsIII. Navigation

CHAPTER 4. MISSION EQUIPMENTSection I. Mission avionics

II. Aircraft survivability equipmentCHAPTER 5. OPERATING LIMITS AND RESTRICTIONS

Section I. GeneralII. System limits

III. Power limitsIV. Loading limits

V. Airspeed limits, maximum and minimumVI. Maneuvering limits

VII. Environmental restrictionsVIII. Other limitations

IX. Required equipment for various conditions of flight

PAGE1-12-12-1

2-222-222-282-352-392-412-532-582-672-702-76

3-13-13-2

3-184-14-14-35-15-15-15-65-75-75-95-9

5-105-10

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CHAPTER 6. WEIGHT/BALANCE AND LOADINGSection I. General

II. Weight and balanceIII. Fuel/oilIV. Center of gravity

V. Cargo loadingCHAPTER 7. PERFORMANCE

Section I. Introduction to performanceCHAPTER 8. NORMAL PROCEDURES

Section I. Mission planningII. Operating procedures and maneuvers

III. Instrument flightIV. Flight characteristics

V. Adverse environmental conditionsVI. Crew duties

CHAPTER 9. EMERGENCY PROCEDURESSection I. Aircraft systems

APPENDIX A. REFERENCESAPPENDIX B. ABBREVIATIONS AND TERMS

INDEX

6-16-16-36-86-86-87-17-18-18-18-1

8-208-218-238-26

9-19-1

A-1B-1

INDEX-1

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CHAPTER 1INTRODUCTION

1-1. GENERAL.

These instructions are for use by the operator(s). Theyapply to the RC-12D aircraft.

1-2. WARNINGS, CAUTIONS, AND NOTES.

Warnings, cautions, and notes are used toemphasize important and critical instructions and areused for the following conditions:

WARNINGAn operating procedure, practice, etc. ,which if not correctly followed, couldresult in personal injury or loss of life.

CAUTIONAn operating procedure, practice, etc., which, if not strictly observed, couldresult in damage to or destruction ofequipment.

NOTEAn operating procedure, condition,etc. , which is essential to highlight.

1-3. DESCRIPTION.

This manual contains the best operatinginstructions and procedures for the RC-12D aircraftunder most circumstances. The observance oflimitations, performance, and weight/balance dataprovided is mandatory. The observance of proceduresis mandatory except when modification is requiredbecause of multiple emergencies, adverse weather,terrain, etc. Your flying experience is recognized, andtherefore basic flight principles are not included. THISMANUAL SHALL BE CARRIED IN THE AIRCRAFT ATALL TIMES.

1-4. APPENDIX A, REFERENCES.

Appendix A is a listing of official publications citedwithin the manual applicable to and available for flightcrews.

1-5. APPENDIX B, ABBREVIATIONS AND TERMS.

Appendix B is a listing of abbreviations and termsused throughout the manual.

1-6. INDEX.

The index lists, in alphabetical order, every titledparagraph, figure, and table contained in this manual.Chapter 7, Performance Data, has an additional indexwithin the chapter.

1-7. ARMY AVIATION SAFETY PROGRAM.

Reports necessary to comply with the safetyprogram are prescribed in AR 385-40.

1-8. DESTRUCTION OF ARMY MATERIEL TOPREVENT ENEMY USE.

For information concerning destruction of Armymateriel to prevent enemy use, refer to TM 750-2441-5.

1-9. FORMS AND RECORDS.

Army aviators flight record and aircraftmaintenance records which are to be used by crewmembers are prescribed in DA PAM 738-751 and TM55-1500-342-23.

1-10. EXPLANATION OF CHANGE SYMBOLS.

Changes, except as noted below, to the text andtables, including new material on added pages, areindicated by a vertical line in the outer margin extendingclose to the entire area of the material affected;exception: pages with emergency markings, whichconsist of black diagonal lines around three edges, mayhave the vertical line or change symbol placed along theinner margins. Symbols show current changes only. Aminiature pointing hand symbol is used to denote achange to an illustration. However, a vertical line in theouter margin, rather than miniature pointing hands, isutilized when there have been extensive changes madeto an illustration. Change symbols are not utilized toindicate changes in the following:

1-1

a. Introductory material.

b. Indexes and tabular data where the changecannot be identified.

c. Blank space resulting from the deletion of text,an illustration or a table.

d. Correction of minor inaccuracies, such asspelling, punctuation, relocation of material, etc., unless correction changes the meaning ofinstructive information and procedures.

1-11. AIRCRAFT DESIGNATION SYSTEM.

The designation system prescribed by AR 70-50 isused in aircraft designations as follows: EXAMPLE RC-12D

R - Modified mission symbol (reconnaissance)

TM 55-1510-219-10C - Basic mission and type symbol (cargo)12 - Design numberD - Series symbol

1-12. USE OF WORDS SHALL, WILL, SHOULD, ANDMAY.

Within this technical manual the word "shall" isused to indicate a mandatory requirement. The word"should" is used to indicate a nonmandatory butpreferred method of accomplishment. The word "may"is used to indicate an acceptable method ofaccomplishment. The word "will" is used to express adeclaration of purpose and may also be used wheresimple futurity is required.

1-13. PLACARD ITEMS.

All placard items (switches, controls, etc. ) areshown throughout this manual in capital letters.

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CHAPTER 2

AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION

Section I. AIRCRAFT

2-1. INTRODUCTION.

The purpose of this chapter is to describe theaircraft and its systems and controls which contribute tothe physical act of operating the aircraft. It does notcontain descriptions of avionics or mission equipment,covered elsewhere in this manual. This chaptercontains descriptive information and does not describeprocedures for operation of the aircraft. Theseprocedures are contained within appropriate chapters inthe manual. This chapter also contains the emergencyequipment installed. This chapter is not designed toprovide instructions on the complete mechanical andelectrical workings of the various systems; therefore,each is described only in enough detail to makecomprehension of that system sufficiently complete toallow for its safe and efficient operation.

2-2. GENERAL.

The RC-12D is a pressurized, low wing, all metal aircraft(figs. 2-1 and 2-2) powered by two PT6A41 turbopropengines, and has all weather capability. Distinguishablefeatures of the aircraft are the slender, streamlinedengine nacelles, an aft rotating boom antenna, missionantennas, antenna pods, wing tip pods, a T-tail and aventral fin below the empennage. The basic mission ofthe aircraft is radio reconnaissance. Cabin entrance ismade through a stair-type door on the left side of thefuselage.

2-3. DIMENSIONS.

Overall aircraft dimensions are shown in Figure 2-3.

2-4. GROUND TURNING RADIUS.

Minimum ground turning radius of the aircraft isshown in Figure 2-4.

2-5. MAXIMUM WEIGHTS.

Maximum takeoff gross weight is 14,200 pounds.Maximum landing weight is 13,500 pounds. Maximumramp weight is 14,290 pounds.

2-6. EXHAUST DANGER AREA.

Danger areas to be avoided by personnel whileaircraft engines are being operated on the ground aredepicted in Figure 2-5. Distance to be maintained withengines operating at idle are shown. Temperature andvelocity of exhaust gases at varying locations aft of theexhaust stacks are shown for maximum power. Thedanger area extends to 40 feet aft of the exhaust stackoutlets. Propeller danger areas are also shown.

2-7. LANDING GEAR SYSTEM.

The landing gear is a retractable, tricycle type,electrically operated by a single DC motor. This motordrives the main landing gear actuators through a gearbox and torque tube arrangement, and also drives achain mechanism which controls the position of the nosegear. Positive down-locks are installed to hold the dragbrace in the extended and locked position. The down-locks are actuated by overtravel of the linear jackscrewsand are held in position by a spring-loaded overcentermechanism. The jackscrew in each actuator holds allthree gears in the UP position, when the gear isretracted. A friction clutch between the gearbox and thetorque shafts protects the motor from electrical overloadin the event of a mechanical malfunction. A 150amperecurrent limiter, located on the DC distribution bus underthe center floorboard, protects against electricaloverload. Rotation of the dipole antenna is controlledthrough the actuation of the right main gear up limitswitch. Gear doors are opened and closed through amechanical linkage connected to the landing gear. Thenose wheel steering mechanism is automaticallycentered and the rudder pedals relieved of the steeringload when the landing gear is retracted. Air-oil typeshock struts, filled with compressed air and hydraulicfluid, are incorporated with the landing gear. Gearretraction or extension time is approximately sixseconds.

a. Landing Gear Control Switch. Landing gearsystem operation is controlled by a manually actuated,wheel-shaped switch placarded LDG GEAR CONTR,UP and DN, on the left subpanel. The control switchand associated relay circuits are protected by a 5-ampere circuit breaker, placarded

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Figure 2-1. General Exterior Arrangement (Sheet 1 of 7)

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2-4

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2-5

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2-6

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2-7

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Figure 2-2. General Interior Arrangement

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Figure 2-3. Principal Dimensions

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Figure 2-4. Ground Turning Radius

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Figure 2-5. Exhaust and Propeller Danger Areas

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TM 55-1510-219-10LANDING GEAR RELAY on the overhead circuit breaker panel (fig. 2-26).

b. Landing Gear Down Position Indicator Lights. Landing gear down position is indicated by three green lights onthe left subpanel, placarded GEAR DOWN. These lights may be checked by operating the ANNUNCIATOR TESTswitch. The circuit is protected by a 5-ampere circuit breaker, placarded LANDING GEAR IND, on the overhead circuitbreaker panel (fig. 2-26).

c. Landing Gear Position Warning Lights. Two red bulbs, wired in parallel and activated by microswitchesindependent of the GEAR DOWN position indicator lights, are positioned inside the clear plastic grip on the landing gearcontrol switch. These lights illuminate whenever the landing gear switch is in either the UP or DN position and the gear isin transit. Both bulbs will also illuminate should either or both power levers be retarded below approximately 79 to 81%NI when the landing gear is not down and locked. To turn the switch lights OFF, during single-engine operation, thepower lever for the inoperative engine must be advanced to a position which is higher than the setting of the warninghorn microswitch. Extending the landing gear will also turn the lights off. Both red lights indicate the same warningconditions, but two are provided for a fail-safe indication in the event one bulb

burns out. The circuit is protected by a 5-ampere circuitbreaker, placarded LANDING GEAR IND, on theoverhead circuit breaker panel (fig. 2-26).

d. Landing Gear Warning Light Test Button. A testbutton, placarded HDL LT TEST, is located on the leftsubpanel. Failure of the landing gear switch toilluminate red, when this test button is pressed, indicatestwo defective bulbs or a circuit fault. The circuit isprotected by a 5-ampere circuit breaker, placardedLANDING GEAR RELAY CONTROL, on the overheadcircuit breaker panel (fig. 2-26).

e. Landing Gear Warning Horn. When eitherpower lever is retarded below approximately 80 ± 1% N

1 with the landing gear out of down and locked positionand the flaps extended beyond 40%, a warning hornlocated in the overhead control panel will soundintermittently. To prevent actuation of the warning hornduring long descents, a pressure differential "Q" switchis connected into the copilot's static air line to preventthe warning horn from sounding at airspeeds greaterthan 140 KIAS. An altitude sensing switch is installed inseries with the "Q" switch to disable the warning horn ataltitudes above approximately 12,500 feet MSL. Thewarning horn is enabled as the aircraft descends throughapproximately 10,500 feet MSL. The warning horncircuit is protected by a 5-ampere circuit breaker,placarded

Figure 2-6. Subpanels

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LANDING GEAR WARN, located on the overheadcircuit breaker panel (fig. 2-26).

f. Landing Gear Warning Horn Test Switch.

The landing gear warning horn may be tested by the testswitch on the right subpanel. The switch, placardedSTALL WARN TEST OFF LDG GEAR WARN TEST,will sound the landing gear warning horn and illuminatethe landing gear position warning lights when moved tothe momentary LDG GEAR WARN TEST position. Thecircuit is protected by a 5-ampere circuit breaker,placarded LANDING GEAR WARN, on the overheadcircuit breaker panel (fig. 2-26).

g. Landing Gear Safety Switches. A safety switchon each main landing gear shock strut controls theoperation of various aircraft systems that function onlyduring flight or only during ground operation. Theseswitches are mechanically actuated whenever the mainlanding gear shock struts are extended (normally aftertakeoff), or compressed (normally after landing). Thesafety switch on the right main landing gear strutactivates the landing gear control circuits, cabinpressurization circuits and the flight hour meter whenthe strut is extended. This switch also activates a down-lock hook, preventing the landing gear from being raisedwhile the aircraft is on the ground. The hook, whichunlocks automatically after takeoff, can be manuallyover-ridden by pressing down on the red button,placarded DN LOCK REL located adjacent to thelanding gear switch. If the over-ride is used and thelanding gear control switch is raised, power will besupplied to the warning horn circuit and the horn willsound. The safety switch on the left main landing gearstrut activates the left and right engine ambient airshutoff valves when the strut is extended.

CAUTIONContinued pumping of handle afterGEAR DOWN position indicator lights(3) are illuminated could damage thedrive mechanism, and preventsubsequent gear retraction.

h. Manual Landing Gear Extension System.Manual landing gear extension is provided through amanually powered system as a backup to the electricallyoperated system. Before manually extending the gear,make certain that the landing gear switch is in the downposition with the LANDING GEAR RELAY circuitbreaker pulled. During a manual landing gearextension, the landing gear motor must be disengagedfrom the landing gear drive mech-

anism. This is accomplished through use of an alternateengage handle located adjacent to the landing gearalternate extension handle. To disengage the landinggear motor, pull the alternate extension handle up andturn it clockwise. When this handle is pulled, thelanding gear motor is disconnected from the system andthe alternate drive system is locked to the gearbox andmotor. With the alternate drive locked in, the landinggear may be manually extended by pumping thealternate extension handle until the three GEAR DOWNposition indicator lamps are illuminated. Refer toChapter 9 for additional information on emergency gearextension procedures.

i. Tires. The aircraft is equipped with dual 22 x 6.75 x 10, 8 ply rated, tubeless, rim inflation tires on themain gear. The nose gear is equipped with a single 22 x6. 75 x 10, 8 ply rated, tubeless, rim inflated tire.

j. Steerable Nose Wheel. The aircraft can bemaneuvered on the ground by the steerable nose wheelsystem. Direct linkage from the rudder pedals (fig. 2-9)to the nose wheel steering linkage allows the nose wheelto be turned 12° to the left of center or 14° to the right.When rudder pedal steering is augmented by the mainwheel braking action, the nose wheel can be deflectedup to 48° either side of center. Shock loads which wouldnormally be transmitted to the rudder pedals areabsorbed by a spring mechanism in the steering linkage.Retraction of the landing gear automatically centers thenose wheel and disengages the steering linkage fromthe rudder pedals.

k. Wheel Brake System. The main wheels areequipped with multiple-disc hydraulic brakes actuated bymaster cylinders attached to the rudder pedals at thepilot's and copilot's position. Braking is permitted fromeither set of rudder pedals. Brake fluid is supplied to thesystem from the reservoir in the nose compartment.The toe brake sections of the rudder pedals areconnected to the master cylinders which actuate thesystem for the or responding wheels. No emergencybrake system is provided.

WARNINGRepeated and excessive application ofbrakes, without allowing sufficientcooling time between applications, willcause loss of braking efficiency, andmay cause brake or wheel failure, tireblowout, or destruction of wheelassembly by fire.

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2-8. PARKING BRAKE.

CAUTIONParking brakes shall not be set duringflight.

Dual parking brake valves are installed below thecockpit floor. Both valves can be closed simultaneouslyby pressing both brake pedals to build up pressure, thenpulling out the handle placarded PARKING BRAKE, onthe left subpanel. Pulling the handle full out sets thecheck valves in the system and any pressure beingapplied by the toe brakes is maintained. Parking brakesare released when the brake handle is pushed in. Theparking brake may be set from either cockpit position.Parking brakes shall not be set during flight.

2-9. ENTRANCE AND EXIT PROVISIONS.

NOTETwo keys are provided in the loosetools and equipment bag. Both keyswill fit the locks on the cabin door,emergency hatch, tailcone accessdoor and the right and left noseavionics doors.

a. Cabin Door.

CAUTIONStructural damage may be caused ifmore than one person is on the cabindoor at a time. The door is weightlimited to less than 300 pounds.

A swing-down door (fig. 2-7), hinged at thebottom, provides a stairway for normal and emergencyentry and exit. Two of the steps are movable and foldflat against the door in the closed position. A step foldsdown over the door sill when the door opens to providea platform (step) for door seal protection. A plasticencased cable provides support for the door in the openposition, a handhold, and a convenience for closing thedoor from inside. A hydraulic damper permits the doorto lower gradually during opening. A rubber seal aroundthe door seals the pressure vessel while the aircraft is inflight. The door locking mechanism is operated byeither of the two mechanically interconnected handles,one inside

and the other outside the door. When either handle isrotated, three rotating-cam-type latches on either side ofthe door capture posts mounted on the cargo door. Inthe closed position, the door becomes an integral part ofthe cargo door. A button adjacent to the door handlemust be depressed before the handle can be rotated toopen the door. A bellows behind the button is inflatedwhen the aircraft is pressurized to prevent accidentalunlatching and/or opening of the door. A small roundwindow just above the second step permits observationof the pressurization safety bellows. A placard adjacentto the window instructs the operator that the safety lockarm is in position around the bellows shaft whichindicates a properly locked door. Pushing the red buttonadjacent to the window will illuminate the inside doormechanism. A CABIN DOOR annunciator light in thecaution/advisory panel will illuminate if the door is notclosed and all latches fully locked.

b. Cargo Door. A swing-up door (fig. 2-7), hingedat the top, provides cabin access for loading cargo orbulky items. After initial opening force is applied, gassprings will completely open the cargo doorautomatically. The door is counterbalanced and willremain in the open position. A door support rod is usedto hold the door in the open position, and to aid inovercoming the pressure of the gas spring assemblieswhen closing the door. Once closed, the gas springsapply a closing force to assist in latching the door. Arubber seal around the door, seals the pressure vesselwhile in flight. The door locking mechanism is operatedonly from inside the aircraft, and is operated by twohandles, one in the bottom forward portion of the doorand the other in the upper aft portion of the door. Whenthe upper aft handle is operated per placard instructions,two rotating cam-type latches on the forward side of thedoor and two on the aft side rotate, capturing postsmounted on the fuselage side of the door opening. Thebottom handle, when operated per placard instructions,actuates four pin lug latches across the bottom of thedoor. A button on the upper aft handle must be pressedbefore the handle can be released to open the door. Alatching lever on the bottom handle must be lifted torelease the handle before the lower latches can beopened. These act as additional aids in preventingaccidental opening or unlatching of the door. The cabinand cargo doors are equipped with dual sensing circuitsto provide the crew remote indication of cabin/cargodoor security. An annunciator light placarded CABINDOOR will illuminate if the cabin or cargo door is openand the BATT switch is ON. If the battery switch isOFF, the annunciator will illuminate only if the cargodoor is not securely closed and latched. The cargo doorsensing circuit receives power from the hot battery bus.

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Figure 2-7. Cabin and Cargo Doors

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CAUTIONInsure the cabin door is closed andlocked. Operating the cargo doorwhile the cabin door is open maydamage the door hinge and adjacentstructure.

(1) Opening cargo door.

CAUTIONAvoid side loading of the gas springsto prevent damage to the mechanism.

1. Handle access door (lower forward cornerof door) -Unfasten and open.

2. Handle Lift hook and move to OPENposition.

3. Handle access door Secure.

4. Handle access door (upper aft corner ofdoor) -Unfasten and open.

5. Handle Press button and lift to OPENposition then latch in place.


7. Door support rod Attach one end to cargodoor ball stud (on forward side of door).

8. Support rod detent pin Check in place.

9. Cabin door sill step Push out on and allowcargo door to swing open. Gas springswill automatically open the door.

10. Door support rod Attach free end to ballstud on forward fuselage door frame.

(2) Closing cargo door.

CAUTIONAvoid side loading of the gas springsto prevent damage to the mechanism.

1. Door support rod Detach from fuselagedoor frame ball stud, then firmly graspfree end of rod

while exerting downward force to overcome the pressureof gas spring assemblies, Then remove support rodfrom door as gas spring assemblies pass over-centerposition.

2. Cargo door Pull closed, using finger holdcavity in fixed cabin door step.

3. Handle access door (upper aft corner ofdoor) -Unfasten and open.

4. Handle Pull handle down until it latches inclosed position.


6. Handle access door (lower forward cornerof door) -Unfasten and open.

7. Handle Move to full forward position.

8. Safety hook Check locked in position bypulling aft on handle.


c. Cabin Emergency Hatch. The cabin emergencyhatch, placarded EXIT-PULL, is located on the rightcabin sidewall just aft of the copilot's seat. The hatchmay be released, from the inside with a pull-downhandle. A flush mounted pull out handle allows thehatch to be released from the outside. The hatch is ofthe non-hinged, plug type which removes completelyfrom the frame when the latches are released. Thehatch can be key locked, from the inside, to preventopening from the outside. The inside handle will unlatchthe hatch whether or not it is locked, by overriding thelocking mechanism. The keylock should be unlockedprior to flight to allow removal of the hatch from theoutside in the event of an emergency. The key remainsin the lock when the hatch is locked and can beremoved only when the hatch is unlocked. The key slotis in the vertical position when the hatch is unlocked.Removal of the key from the lock before flight assuresthe pilot that the hatch can be removed from the outsideif necessary.

2-10. CABIN DOOR CAUTION LIGHT.

As a safety precaution, two illuminated MASTERCAUTION lights on the glare shield and a steadilyilluminated CABIN DOOR yellow caution light on theannunciator panel indicates the cabin door is not closedand locked. This circuit is protected by 5-ampere circuitbreakers placarded ANN PWR and ANN IND, located onthe overhead circuit breaker panel (fig. 2-26).

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2-11. WINDOWS.

a. Cockpit Windows. The pilot and copilot haveside windows, a windshield and storm windows, whichprovide visibility from the cockpit. The storm windowsmay be opened on the ground or during unpressurizedflight.

b. Cabin Windows. The outer cabin windows, of two-ply construction, are the pressure type and are integralparts of the pressure vessel. All cabin windows arepainted over except for the two small aft windows(movable curtains are provided to cover these windowswhen it is desited to seal out light).

2-12. SEATS.

a. Pilot and Copilot Seats. The pilot and copilotseats (fig. 2-8) are separated from the cabin bymovable curtains. The controls for vertical heightadjustment and fore and aft travel are located undereach seat. The fore and aft adjustment handle islocated beneath the bottom front inboard corner of eachseat. Pulling up on the handle allows the seat to movefore or aft. The height adjustment handle is locatedbeneath the bottom front outboard corner of

each seat. Pulling up on the handle, allows the seat tomove up and down. Both seats have adjustableheadrests and armrests which will raise and lower foraccess to the cockpit. Handholds on either side of theoverhead panels and a fold-away protective pedestalstep are provided for pilot and copilot entry into thecockpit (fig. 2-9). For the storage of maps and theoperator's manual, pilot and copilot seats have aninboard-slanted, expandable pocket affixed to the lowerportion of the seat back. Pocket openings are heldclosed by shock cord tension.

b. Pilot and Copilot Seat Belts and ShoulderHarnesses. Each pilot and copilot seat is equipped witha lap-type seat belt and shoulder harness connected toan inertia reel. The shoulder harness belt is of the "Y"configuration with the single strap being contained in aninertia reel attached to the base of the seatback. Thetwo straps are worn with one strap over each shoulderand fastened by metal loops into the seat belt buckle.The spring loading at the inertia reel keeps the harnesssnug but will allow normal movement required duringflight operations.

The inertia reel is designed with a locking device thatwill secure the harness in the event of sudden forwardmovement or an impact action.

Figure 2-8. Pilot and Copilot Seats

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Figure 2-9. Cockpit

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Figure 2-10. PT6A-41 Engine (Sheet 1 of 2)

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Figure 2-10. PT6A-41 Engine (Sheet 2 of 2)

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Section II: EMERGENCY EQUIPMENT

2-13. DESCRIPTION.

The equipment covered in this section includes allemergency equipment, except that which forms part of acomplete system. For example, landing gear system,etc. Chapter 9 describes the operation of emergencyexits and location of all emergency equipment.

2-14. FIRST AID KITS.

Four first aid kits are included in the survival kit.

2-15. HAND-OPERATED FIRE EXTINGUISHER.

WARNINGRepeated or prolonged exposure tohigh concentrations ofmonobromotrifluo-

romethane (CF3Br) or decompositionproducts should be avoided. Theliquid shall not be allowed to comeinto contact with the skin, as it maycause frost bite or low temperatureburns because of its very low boilingpoint.

One hand-operated fire extinguisher is mountedbelow the pilot's seat and a second extinguisher, locatedon the left cabin sidewall, aft of the cabin door. Theyare of the monobromotrifluoromethane (CF3Br) type.The extinguisher is charged to a pressure of 150 to 170PSI and emits a forceful stream. Use an extinguisherwith care within the limited area of the cabin to avoidsevere splashing.

NOTEEngine fire extinguisher systems aredescribed in Section III.

Section III. ENGINES AND RELATED SYSTEMS

2-16. DESCRIPTION.

The aircraft is powered by two PT6A-41 turbopropengines (fig. 2-10). The engine has a three stage axial,single stage centrifugal compressor, driven by a singlestage reaction turbine. The power turbine, a two stagereaction turbine, counter-rotating with the compressorturbine, drives the output shaft. Both the compressorturbine and the power turbine are located in theapproximate center of the engine with their shaftsextending in opposite directions. Being a reverse flowengine, the ram air supply enters the lower portion of thenacelle and is drawn in through the aft protectivescreens. The air is then routed into the compressor.After it is compressed, it is forced into the annularcombustion chamber, and mixed with fuel that issprayed in through 14 nozzles mounted around the gasgenerator case. A capacitance discharge ignition unitand two spark igniter plugs are used to start combustion.After combustion, the exhaust passes through thecompressor turbine and two stages of power turbine andis routed through two exhaust ports near the front of theengine. A pneumatic fuel control system schedules fuelflow to maintain the power set by the gas generatorpower lever. The accessory drive at the aft end of theengine provides power to drive the fuel pumps,

fuel control, the oil pumps, the refrigerant compressor(right engine), the starter/generator, and the turbinetachometer transmitter. The reduction gearbox forwardof the power turbine provides gearing for the propellerand drives the propeller tachometer transmitter, thepropeller overspeed governor, and the propellergovernor. A torque limiter is incorporated on the frontcase of the propeller reduction gearbox to maintaindeveloped engine torque within design limits.

2-17. ENGINE COMPARTMENT COOLING.

The forward engine compartment including theaccessory section is cooled by air entering around theexhaust stack cutouts, the gap between the propellerspinner and forward cowling, and exhausting throughducts in the upper and lower aft cowling.

2-18. AIR INDUCTION SYSTEMS GENERAL.

Each engine and oil cooler receives ram airducted from an air scoop located within the lowersection of the forward nacelle. Special components ofthe engine induction system protect the power plantfrom icing and foreign object damage.

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Figure 2-11. Pedestal

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2-19. FOREIGN OBJECT DAMAGE CONTROL.

The engine has an integral air inlet screendesigned to obstruct objects large enough to damage thecompressor.

2-20. ENGINE ICE PROTECTION SYSTEMS.

a. Inertial Separator.

CAUTIONAfter the ice vanes have beenmanually extended, they may bemechanically actuated only. Noelectrical extension or retractionshall be attempted as damage tothe actuator may result. Linkage inthe nacelle area must be reset priorto operation of the electric system.

An inertial separation system is built into eachengine air inlet to prevent moisture particles from enteringthe engine inlet plenum under icing conditions. Amovable vane and a bypass door are lowered into theairstream when operating in visible moisture at 5'C orcolder, by energizing electrical actuators with theswitches, placarded ICE VANE RETRACT EXTEND,located on the overhead control panel. A mechanicalbackup system is provided, and is actuated by pulling theT-handles just below the pilot's subpanel placarded ICEVANE #1 ENG #2 ENG. Decrease airspeed to 160 knotsor less to reduce forces for manual extension. Normalairspeed may then be resumed.

(1) The vane deflects the ram airstream slightlydownward to introduce a sudden turn in the airstream tothe engine, causing the moisture particles to continue onundeflected, because of their greater momentum, and tobe discharged overboard.

(2) While in the icing flight mode, the extendedposition of the vane and bypass door is indicated by greenannunciator lights, #1 VANE EXT and #2 VANE EXT.

(3) In the non-ice protection mode, the vaneand bypass door are retracted out of the airstream byplacing the ice vane switches in the RETRACT position.The green annunciator lights will extinguish. Retractionshould be accomplished at 15°C and above to assureadequate oil cooling. The vanes should be eitherextended or retracted; there are no intermediate positions.

(4) If for any reason the vane does not attainthe selected position within approximately 15 seconds, ayellow #1 VANE FAIL or #2 VANE FAIL light illuminateson the caution/advisory panel. In this event, the manualbackup system should be used. When the vane issuccessfully positioned with the manual system, theyellow annunciator lights will extinguish.

b. Engine Air Inlet Deice System.

(1) Description. Hot engine exhaust gas isutilized for heating the air inlet lips to prevent theformation of ice. Hot exhaust gas is picked up insideeach engine exhaust stack and carried by plumbing to theinlet lip. The gas flows through the inside of the lip to thebottom where it is allowed to escape.

(2) Engine air inlet deice system switches.(Provisions only system not installed. ) Two switchesplacarded ENG INLET LIP HEAT #1, #2 OFF/ ON,located on the overhead control panel, operate solenoidvalves in the exhaust system of each engine. Thesevalves control the flow of hot exhaust gases to the inlet airlip assemblies.

(3) Fuel heater. A oil-to-fuel heat exchanger,located on the engine accessory case, operatescontinuously and automatically to heat the fuel sufficientlyto prevent ice from collecting in the fuel control unit.Each pneumatic fuel control line is protected against ice.Power is supplied to each fuel control air line jacketheater by two switches actuated by moving the conditionlevers in the pedestal out of the fuel cutoff range. Fuelcontrol heat is automatically turned on for all engineoperations.

2-21. ENGINE FUEL CONTROL SYSTEM.

a. Description. The basic engine fuel systemconsists of an engine driven fuel pump, a fuel control unit,a fuel flow divider, a dual fuel manifold and fourteen fuelnozzles.

b. Fuel Control Unit. One fuel control unit ismounted on the accessory case of each engine. This unitis a hydro-mechanical metering device which determinesthe proper fuel schedule for the engine to produce theamount of power requested by the relative position of itspower lever. The control of developed engine power isaccomplished by adjusting the engine compressor turbine(N1) speed. N speed controlled by varying the amount offuel injected into the combustion chamber through thefuel nozzles. Engine shutdown is accomplished bymoving the appropriate condition lever to the full aft,FUEL CUT-OFF position, which shuts off the fuel supply.

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2-22. POWER LEVERS.

CAUTIONMoving the power levers intoreverse range without the enginesrunning may result in damage tothe reverse linkage mechanism.

Two power levers are located on the controlpedestal (fig. 2-11). These levers regulate power in thereverse, idle, and forward range, and operate so thatforward movement increases engine power. Powercontrol is accomplished through adjustment of the N1speed governor in the fuel control unit. Power isincreased when Ni RPM is increased. The power leversalso control propeller reverse pitch. Distinct movement(pulling up and then aft on the power lever) by the pilot isrequired for reverse thrust. Placarding beside the levertravel slots reads POWER. Upper lever travel range isdesignated INCR (increase), supplemented by an arrowpointing forward. Lower travel range is marked IDLE,LIFT and REVERSE. A placard below the lever slotsreads: CAUTION REVERSE ONLY WITH ENGINESRUNNING.

2-23. CONDITION LEVERS.

Two condition levers are located on the controlpedestal (fig. 2-11). Each lever starts and stops the fuelsupply, and controls the idle speed for its engine. Thelevers have three placarded positions: FUEL CUTOFF,LO IDLE, and HIGH IDLE. In the FUEL CUTOFFposition, the condition lever controls the cutoff function ofits engine-mounted fuel control unit. From LO IDLE toHIGH IDLE, they control the governors of the fuel controlunits to establish minimum fuel flow levels. LO IDLEposition sets the fuel flow rate to attain 52 to 55% (at sealevel) minimum NI and HIGH IDLE position sets the rateto attain 70% minimum NI. The power lever for thecorresponding engine can select NI from the respectiveidle setting to maximum power. An increase in low idle Niwill be experienced at high field elevation.

2-24. FRICTION LOCK KNOBS.

Four friction lock knobs, placarded FRICTIONLOCK, are provided on the control pedestal (fig. 211) toadjust friction drag against the engine power, propellerRPM, and fuel condition levers. These knobs prevent thelevers from creeping. When rotated clockwise, each knobincreases the friction that opposes movement of theaffected lever. Counterclockwise movement of the knobdecreases friction.

2-25. ENGINE FIRE DETECTION SYSTEM.

a. Description. A flame surveillance system isinstalled on each engine to detect external engine fire andprovide alarm to the pilot. Both nacelles are monitored,each having a control amplifier and three detectors.Electrical wiring connects all sensors and controlamplifiers to DC power and to the cockpit visual alarmunits. In each nacelle, one detector monitors the forwardnacelle, a second monitors the upper accessory area, anda third the lower accessory area. Fire emits an infraredradiation that will be sensed by the detector whichmonitors the area of origin. Radiation exposure activatesthe relay circuit of a control amplifier which causes signalpower to be sent to cockpit warning systems. Anactivated surveillance system will return to the standbystate after the fire is out. The system includes afunctional test switch and has circuit protection throughthe FIRE DETR circuit breaker. Warning of internalnacelle fire is provided as follows: the red MASTERWARNING lights on the glareshield illuminateaccompanied by the illumination of a red warning light inthe appropriate fire control T-handle placarded FIREPULL (fig. 2-28). Fire detector circuits are protected by asingle 5-ampere circuit breaker, placarded FIRE DETR,located on the overhead circuit breaker panel (fig. 2-26).

b. Fire Detection System Test Switch. Onerotary switch placarded FIRE PROTECTION TEST on thecopilot's subpanel is provided to test the engine firedetection system. Before checkout, battery power mustbe on and the FIRE DETR circuit breaker must be closed.Switch position DETR 1, checks the area forward of theair intake of each nacelle, including circuits to the cockpitalarm and indication devices. Switch position DETR 2,checks the circuits for the upper accessory compartmentof each nacelle. Switch position DETR 3, checks thecircuits for the lower accessory compartment of eachnacelle. Each numbered switch position will initiate thecockpit indications previously described.

c. Erroneous Fire Detection SystemIndications. During ground test of the engine firedetection system, an erroneous indication of system faultmy be encountered if an engine cowling is not closedproperly, or if the aircraft is headed toward a strongexternal light source. In this circumstance, change theaircraft heading to enable a valid system check.

2-26. ENGINE FIRE EXTINGUISHER SYSTEM.

a. Description. The fire extinguisher systemutilizes an explosive squib and valve which, when

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opened, allows the distribution of the pressurizedextinguishing agent through a plumbing network of spraynozzles strategically located in the fire zones of theengines.

b. Fire Pull Handles. The fire control handles,which are used to arm the extinguisher system arecentrally located on the pilot's instrument panel (fig. 2-28), immediately below the glareshield. These controlsreceive power from the hot battery bus. The fire detectionsystem will indicate an engine fire by illuminating themaster fault warning light on the pilot's and copilot'sglareshield and the respective #1 or #2 FIRE PULL lightsin the fire control T-handles. Pulling the fire control T-handle will electrically arm the extinguisher system andclose the fuel firewall shutoff valve for that particularengine. This will cause the red light in the PUSH TOEXTINGUISH switch and the respective red #1 and #2FUEL PRESS light in the warning annunciator panel toilluminate. Pressing the lens of the PUSH TOEXTINGUISH switch (after lifting one side of its spring-loaded clear plastic guard) will fire the squib, expelling allthe agent in the cylinder at one time. The respectiveyellow caution light, #1 or #2 EXTGH DISCH on thecaution/advisory annunciator panel will illuminate andremain illuminated until the squib is replaced.

c. Fire Extinguisher System Test Switch. A rotarytest switch, placarded FIRE PROTECTION TEST, islocated on the copilot's subpanel. The test functions,placarded EXTGH #1 #2, are arranged on the left side ofthe switch and provide a test of the pyrotechnic cartridgecircuitry. During Before Exterior Check, the pilot shouldrotate the test switch through the two positions and verifythe illumination of the green SQUIB OK light on thePUSH TO EXTINGUISH switch and the correspondingyellow #1 or #2 EXTGH DISCH light on thecaution/advisory annunciator panel.

d. Fire Extinguishing System Supply CylinderGages. A gage, calibrated in PSI, is mounted on eachsupply cylinder for determining the level of charge andshould be checked during preflight (table 2-1).

2-27. OIL SUPPLY SYSTEM.

a. The engine oil tank is integral with the air-inletcasting located forward of the accessory gearbox. Oil forpropeller operation, lubrication of the reduction gearboxand engine bearings is supplied by an external line fromthe high pressure pump. Two scavenge lines return oil tothe tank from the propeller reduction gearbox. A non-congealing external oil cooler keeps the engine oiltemperature within the operating limits. The capacity ofeach engine oil tank is 2. 3 U. S. gallons. The totalsystem capacity for each engine, which includes the oiltank, oil cooler, lines, etc. , is 3. 5 U. S. gallons. Theoil level is indicated by a dipstick attached to the oil fillercap. Oil grade, specification and servicing points, aredescribed in Section XII, Servicing.

b. The oil system of each engine is coupled to aheat exchanger unit (radiator) of fin-and-tube design.These exchanger units are the only airframe mounted partof the oil system and are attached to the nacelles belowthe engine air intake. Each heat exchanger incorporatesa thermal bypass which assists in maintaining oil at theproper temperature range for engine operation.

2-28. TORQUE LIMITER

A torque limiter is installed in the torquemeterpressure transmitter boss on the front case of thepropeller reduction gearbox. The limiter incorporates asealed bellows, connected directly to torquemeter oilpressure, which works against an externally adjustabletorque-limit spring. Torquemeter oil pressure is sensed atthe inside of the bellows assembly, and during normaloperating conditions the torquemeter oil pressure is nothigh enough to compress the torque-limit spring. Whenhigh torque pressure is sensed, the bellows will stretchand compress the spring assembly. This will permit alimited amount of governing air pressure to be bled fromthe fuel control unit and will continue to do so until enginespeed is reduced, resulting in a proportional reduction inengine torque pressure.

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2-29. ENGINE CHIP DETECTION SYSTEM.

A magnetic chip detector is installed in the bottomof each engine nose gearbox to warn the pilot of oilcontamination and possible engine failure. The sensor isan electrically insulated gap immersed in the oil,functioning as a normally-open switch. If a large metalchip or a mass of small particles bridge the detector gap,a circuit is completed, sending a signal to illuminate a redannunciator panel indicator light placarded #1 or #2 CHIPDETR and the MASTER WARNING lights. Chip detectorcircuits are protected by two 5-ampere circuit breakers,placarded CHIP DETR #1 and #2 on the overhead circuitbreaker panel (fig. 2-26).

2-30. ENGINE IGNITION SYSTEM.

a. The basic ignition system consists of a solid stateignition exciter unit, two igniter plugs, two shielded ignitioncables, pilot controlled IGNITION AND ENGINE STARTswitches and the ENG AUTO IGN switch. Placing anIGNITION AND ENGINE START switch to ON (forward)will cause the respective igniter plugs to spark, ignitingthe fuel/air mixture sprayed into the combustion chamberby the fuel nozzles. The ignition system is activated forground and air starts, but is switched off after combustionlight up.

b. One three-position toggle switch for each engine,located on the overhead control panel, will initiate startermotoring and ignition in the ON/ ENG START position, orwill motor the engine in the STARTER ONLY (aft) position(fig. 2-18). The switches are placarded #1 ENG STARTor #2 ENG START to designate the appropriate engine.The ON switch position completes the starter circuit forengine rotation, energizes the igniter plugs for fuelcombustion, and activates the IGN ON light on theannunciator panel. At center position the switch is OFF.Two 5-ampere circuit breakers on the overhead circuitbreaker panel, placarded IGNITOR CONTR #1 and #2,protect ignition circuits. Two 5ampere circuit breakers onthe overhead circuit breaker panel, placarded STARTCONTR #1 and #2, protect starter control circuits (fig. 2-26).

2-31. AUTO IGNITION SYSTEM.

If armed, the auto ignition system automaticallyprovides combustion re-ignition of either engine shouldaccidental flameout occur. The system is not essential tonormal engine operation, but is used to reduce thepossibility of power loss due to icing or other conditions.Each engine has a separate auto ignition control switchand a green indicator light placarded #1 or #2 IGN ON, onthe annunciator panel. Auto ignition is accomplished byenergizing the two igniter plugs in each engine.

NOTEThe system should be turned OFFduring extended ground operationto prolong the life of the igniterplugs.

a. Auto Ignition Switches. Two switches placardedENG AUTO IGN #1 or #2, control the auto ignitionsystems. The ARM position initiates a readiness modefor the auto ignition system of the corresponding engine.The OFF position disarms the system. Each switch isprotected by a corresponding START CONTR #1 or #2 5-ampere circuit breaker on the overhead circuit breakerpanel (fig. 2-26).

b. Auto Ignition Lights. If an armed auto ignitionsystem changes from a ready condition to an operatingcondition (energizing the igniter plugs in the engine) acorresponding green annunciator panel light willilluminate. The annunciator panel light is placarded #1 or#2 IGN ON and indicates that the igniters are energized.The auto ignition system is triggered from a readycondition to an operating condition when engine torquedrops below approximately 20%. Therefore, when anauto ignition system is armed, the igniters will beenergized continuously during the time when an engine isoperating at a level below approximately 20% torque.The auto ignition lights are protected by 5-ampereIGNITOR CONTR #1 or #2 circuit breakers, located onthe overhead circuit breaker panel (fig. 2-26).

2-32. ENGINE STARTER-GENERATORS.

One start-generator is mounted on each engineaccessory drive section. Each is able to function either asa starter or as a generator. In the starter function, 28volts DC is required to power rotation. In the generatorfunction, each unit is capable of 400 amperes DC output.Each starter circuit is protected by a 5-ampere circuitbreaker placarded START CONTR, located on theoverhead circuit breaker panel. An automatic starter cut-off is installed in each starter-generator to provideautomatic termination of the start cycle at approximately40% NI. This rotational speed is sensed by a magneticsensor internal to the starter-generator. For additionalinformation on the starter-generator system refer toSection IX.

2-33. ENGINE INSTRUMENTS.

The engine instruments are vertically mountednear the center of the instrument panel (fig. 2-28).

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a. Turbine Gas Temperature Indicators. Two TGTgages on the instrument panel (fig. 2-28) are calibrated indegrees Celsius. Each gage is connected tothermocouple probes located in the hot gases betweenthe turbine wheels. The gages register the temperaturepresent between the compressor turbine and powerturbine for the corresponding engine.

b. Engine Torquemeters. Two torquemeters on theinstrument panel (fig. 2-28) indicate torque applied to thepropeller shafts of the respective engines. Each gageshows torque in percent of maximum using 2 percentgraduations and is actuated by an electrical signal from apressure sensing system located in the respectivepropeller reduction gear case. Torquemeters areprotected by individual 0. 5ampere circuit breakersplacarded TORQUEMETER #1 or #2 on the overheadcircuit breaker panel (fig. 2-26).

c. Turbine Tachometers. Two tachometers on theinstrument panel (fig. 2-28) register compressor turbineRPM (N1) for the respective engine. These indicatorsregister turbine RPM as a percentage of maximum gasgenerator RPM. Each instrument is slaved to atachometer generator attached to the respective engine.

d. Oil Pressure/Oil Temperature Indicators. Twogages on the instrument panel (fig. 2-28) panel registeroil pressure in PSI and oil temperature in °C. Oil pressureis taken from the delivery side of the main oil pressurepump. Oil temperature is transmitted by a thermal sensorunit which senses the temperature of the oil as it leavesthe delivery side of the oil pressure pump. Each gage isconnected to pressure transmitters installed on therespective engine. Both instruments are protected by5ampere circuit breakers, placarded OIL PRESS and OILTEMP #1 or #2, on the overhead circuit breaker panel(fig. 2-26).

e. Fuel Flow Indicators. Two gages on theinstrument panel (fig. 2-28) register the rate of flow forconsumed fuel as measured by sensing units coupled intothe fuel supply lines of the respective engines. The fuelflow indicators are calibrated in increments of hundreds ofpounds per hour. Both circuits are protected by 1/2ampere circuit breakers placarded FUEL FLOW #1 or #2,on the overhead circuit breaker panel (fig. 2-26).

Section IV. FUEL SYSTEM

2-34. FUEL SUPPLY SYSTEM.

The engine fuel supply system (fig. 2-12) consistsof two identical systems sharing a common fuelmanagement panel and fuel crossfeed plumbing. Eachfuel system consists of five interconnected wing tanks, anacelle tank, an auxiliary inboard fuel tank. A fueltransfer pump is located within each auxiliary tank.Additionally, the system has an engine-driven boostpump, a standby fuel pump located within each nacelletank, a fuel heater (engine oilto-fuel heat exchanger unit),a tank vent system, a tank vent heating system andinterconnecting wiring and plumbing. Refer to Section IVfor fuel grades and specifications. Fuel tank capacity isshown in Table 2-2.

a. Engine Driven Boost Pumps.

CAUTIONEngine operation using only theenginedriven primary (highpressure) fuel pump withoutstandby pump or engine-drivenboost pump fuel pressure is limitedto 10 cumulative hours. Thiscondition is indicated byillumination of either #1 or #2 FUELPRESS lights and the simultaneousillumination of both MASTER

WARNING lights. Refer to Chapter9. All time in this category shall beentered on DA Form 2408-13 for theattention of maintenancepersonnel.

A gear driven boost pump, mounted on eachengine supplies fuel under pressure to the inlet of theengine-driven primary high-pressure pump for enginestarting and all normal operations. Either the engine-driven boost pump or standby pump is capable ofsupplying sufficient pressure to the enginedriven primaryhigh-pressure pump and thus maintain normal engineoperation.

b. Standby Fuel Pumps. A submerged, electrically-operated standby fuel pump, located within each nacelletank, serves as a backup unit for the engine-driven boostpump. The standby pumps are switched off duringnormal system operations. A standby fuel pump will beoperated during crossfeed to pump fuel from one systemto the opposite engine. The correct pump is automaticallyselected when the CROSSFEED switch is activated.Each standby fuel pump has an inertia switch included inthe power supply circuit. When subjected to a 5 to 6 Gshock loading, as in a crash situation, the inertia switchwill remove electrical power from the standby fuel pumps.The standby fuel pumps are protected

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by two 10-ampere circuit breakers placarded STANDBYPUMP #1 or #2, located on the overhead circuit breakerpanel (fig. 2-26), and four 5ampere circuit breakers (2each in parallel) on the hot battery bus.

c. Fuel Transfer Pumps. The auxiliary tank fueltransfer system automatically transfers the fuel from theauxiliary tank to the nacelle tank without pilot action.Motive flow to a jet pump mounted in the auxiliary tanksump is obtained from the engine fuel plumbing systemdownstream from the engine driven boost pump androuted through the transfer control motive flow valve.The motive flow valve is energized to the open positionby the control system to transfer auxiliary fuel to thenacelle tank to be consumed by the engine during theinitial portion of the flight. When an engine is started,pressure at the engine driven boost pump closes apressure switch which, after a 30 to 50 second time delayto avoid depletion of fuel pressure during starting,energizes the motive flow valve. When auxiliary fuel isdepleted, a low level float switch de-energizes the motiveflow valve after a 30 to 60 second time delay provided toprevent cycling of the motive flow valve due to sloshingfuel. In the event of a failure of the motive flow valve orthe associated control circuitry, the loss of motive flowpressure when there is still fuel remaining in the auxiliaryfuel tank is sensed by a pressure switch and float switch,respectively, which illuminates a light placarded #1 or #2NO FUEL XFR on the annunciator panel. During enginestart, the pilot should note that the NO FUEL XFR lightsextinguish 30 to 50 seconds after engine start. The NOFUEL XFR lights will not illuminate if auxiliary tanks areempty. A manual override is incorporated as a backup forthe automatic transfer system. This is initiated by placingthe AUX TRANSFER switch, located on the fuelmanagement panel (fig. 2-13) to the OVERRIDE position.This will energize the transfer control motive flow valve.The transfer systems are protected by 5-ampere circuit

breakers placarded AUXILIARY TRANSIEK #1 or #2,located on the overhead circuit breaker panel (fig. 2-26).

NOTEIn turbulence or during maneuvers,the NO FUEL XFR lights maymomentarily illuminate after theauxiliary fuel has completedtransfer.

d. Fuel Gaging System. The total fuel quantity inthe left or right main system or left or right auxiliary tankis measured by a capacitance type fuel gaging system.Two fuel gages, one for the left and one for the right fuelsystem, read fuel quantity in pounds. Refer to Section XIIfor fuel capacities and weights. A maximum of 3% errormay be encountered in each system. However, thesystem is compensated for fuel density changes due totemperature excursions. In addition to the fuel gages,yellow No. 1 or No. 2 NAC LOW lights on the caution/advisory annunciator panel illuminate when there isapproximately 20 minutes of fuel per engine remaining(on standard day, at Sea Level, Maximum Cruise Powerconsumption rate). The fuel gaging system is protectedby individual 5 ampere circuit breakers placarded QTYIND and QTY WARN #1 or #2, located on the overheadcircuit breaker panel (fig. 226). A mechanical spiral floatgage (fig. 2-13) is installed in the auxiliary fuel tank toprovide an indication of fuel level when servicing the tank.The gage is installed on the auxiliary fuel tank cover,adjacent to the filler neck. A small sight window in theupper wing skin permits observation of the gage.

e. Fuel Management Panel. The fuel managementpanel (fig. 2-13) is located on the cockpit overheadbetween the pilot and copilot. It contains the fuel gages,standby fuel pump switches, the crossfeed valve switchand a fuel gaging system control switch and transfercontrol switches are also installed.

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Table 2-2. Fuel Quantity Data

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Figure 2-12. Fuel System Schematic

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Figure 2-13. Fuel Management Panel and Auxiliary Tank Mechanical Gage

(1) Fuel gaging system control switch. Aswitch on the fuel management panel (fig. 2-13)placarded FUEL QUANTITY, MAIN AUXILIARY, controlsthe fuel gaging system. When in the MAIN position thefuel gages read the total fuel quantity in the left and rightwing fuel system. When in the AUXILIARY position thefuel gages read the fuel quantity in the left and rightauxiliary tanks only.

(2) Standby fuel pump switches. Twoswitches, placarded STANDBY PUMP ON located on thefuel management panel (fig. 2-13) control a submergedfuel pump located in the corresponding nacelle tank.During normal aircraft operation both switches are off solong as the engine-driven boost pumps function andduring crossfeed operation. The loss of fuel pressure,due to failure of an engine driven boost pump willilluminate the MASTER WARNING lights on theglareshield and will illuminate the #1 FUEL PRESS or #2FUEL PRESS on the warning annunciator panel. TurningON the STANDBY PUMP will extinguish the FUELPRESS lights. The MASTER WARNING lights must bemanually cleared.

NOTEBoth standby pump switches shallbe off during crossfeed operation.

(3) Fuel transfer control switches. Twoswitches on the fuel management panel (fig. 2-13),placarded AUX TRANSFER OVERRIDE AUTO controloperation of the fuel transfer pumps. During normaloperation both switches are in AUTO which allows thesystem to be automatically actuated by fuel flow to theengine. If either transfer system fails to operate, the faultcondition is indicated by two illuminated MASTERCAUTION lights on the glareshield and a steadilyillumihated yellow #1 or #2 NO FUEL XFR light on thecaution annunciator panel.

(4) Fuel crossfeed switch. The fuel crossfeedvalve is controlled by a 3-position switch (fig, 2-13),located on the fuel management panel, placardedCROSSFEED OFF. Under normal flight conditions theswitch is left in the OFF position. During emergencysingle engine operation, it may become necessary tosupply fuel to the operative engine from the fuel systemon the opposite side. The crossfeed system (fig. 2-15) isplacarded for fuel selection with a simplified diagram onthe overhead fuel control panel. Place the standby fuelpump switches in the off position when crossfeeding. Alever lock switch, placarded CROSSFEED, is moved fromthe center OFF position to the left or to the right,depending on direction of fuel flow. This opens thecrossfeed valve and energizes the standby

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Figure 2-14. Gravity Feed Fuel Flow

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Figure 2-15. Crossfeed Fuel Flow

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pump on the side from which crossfeed is desired. Duringcrossfeed operation with firewall fuel valve closed,auxiliary tank fuel will not crossfeed. When the crossfeedmode is energized, a green FUEL CROSSFEED light onthe caution/advisory panel will illuminate. Crossfeedsystem operation is described in Chapter 9. Thecrossfeed valve is protected by a 5-ampere circuit breakerplacarded CROSSFEED VALVE located on the overheadcircuit breaker panel (fig. 2-26).

f. Firewall Shutoff Valves.

CAUTIONDo not use the fuel firewall shutoffvalve to shut down an engine,except in an emergency. Theengine-driven highpressure fuelpump obtains essential lubricationfrom fuel flow. When an engine isoperating, this pump may beseverely damaged (while cavitating)if the firewall valve is closed beforethe condition lever is moved to theFUEL CUTOFF position.

The fuel system incorporates a fuel line shutoffvalve mounted on each engine firewall. The firewallshutoff valves close automatically when the fireextinguisher T-handles on the instrument panel are pulledout. The firewall shutoff valves receive electrical powerfrom the main buses and also from the hot battery buswhich is connected directly to the battery. The valves areprotected by circuit breakers placarded FIREWALLVALVE #1 or #2 on the overhead circuit breaker panel(fig. 2-26), and FIREWALL SHUTOFF #1 or #2 on thehot battery bus circuit breaker board.

g. Fuel Sump Drains. A sump drain wrench isprovided in the aircraft loose tools to simplify draining asmall amount of fuel from the sump drain. There are fivesump drains and one filter drain in each wing (Table 2-3).

An additional drain for the extended range fuel systemline extends through the bottom of the wing center sectionadjacent to the fuselage. Anytime the extended rangesystem is in use, a part of the preflight inspection wouldconsist of draining a small amount of fuel from this drainto check for fuel contamination. Whenever the extendedrange system is removed from the aircraft and the fuelline is capped off in the fuselage, the remaining fuel in theline shall be drained.

h. Fuel Drain Collector System. Each engine isprovided with a fuel drain collector system to return fueldumped from the engine during clearing and shutdownoperations back into its respective nacelle tank. Thesystem draws power from the #4 feeder bus. Fueltransfer is completely automatic. Fuel from the engineflow divider drains into a collector tank mounted below theaft engine accessory section. An internal float switchactuates an electric scavenger pump which delivers thefuel to the fuel purge line just aft of the fuel purge shutoffvalve. A check valve in the line prevents the backflow offuel during engine purging. The circuit breaker for bothpumps is located in the fuel section of the overheadcircuit breaker panel; placarded SCAVENGER PUMP. Avent line, plumbed from the top of the collector tank, isrouted through an inline flame arrestor and thendownward to a drain manifold on the underside of thenacelle.

i. Fuel Vent System. Each fuel system is ventedthrough two ram vents located on the underside of thewing adjacent to the nacelle. To prevent icing of the ventsystem, one vent is recessed into the wing and thebackup vent protrudes out from the wing and contains aheating element. The vent line at the nacelle contains aninline flame arrestor.

j. Engine Oil-to-Fuel Heat Exchanger. An engineoil-to-fuel heat exchanger, located on each engineaccessory case, operates continuously and automaticallyto heat the fuel delivered to the engine sufficiently toprevent the freezing of any water which it might contain.The temperature of the delivered fuel is thermostaticallyregulated to remain between 21°C and 32°C.

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Table 2-3. Fuel Sump Drain Locations

TM 55-1510-219-10

2-35. FUEL SYSTEM MANAGEMENT.

a. Fuel Transfer System. When the auxiliary tanksare filled, they will be used first. During transfer ofauxiliary fuel, which is automatically controlled, thenacelle tanks are maintained full. A check valve in thegravity feed line from the outboard wing prevents reversefuel flow. Normal gravity transfer of the main wing fuelinto the nacelle tanks will begin when auxiliary fuel isexhausted. The system will gravity feed fuel only to itsrespective nacelle tank, i. e. left or right. Fuel will notgravity feed through the crossfeed system.

b. Operation With Failed Engine-Driven Boost Pumpor Standby Pump. Two pumps in each fuel systemprovide inlet head pressure to the engine driven primaryhigh-pressure fuel pump. If crossfeed is used, a thirdpump, the standby fuel pump from the opposite system,

will supply the required pressure. Operation under thiscondition will result in an unbalanced fuel load as fuelfrom one system will be supplied to both engines while allfuel from the system with the failed engine driven andstandby boost pumps will remain unused. A triple failure,which is highly unlikely, would result in the engine drivenprimary pump operating without inlet head pressure.Should this situation occur, the affected engine cancontinue to operate from its own fuel supply on its engine-driven primary high-pressure fuel pump.

2-36. FERRY FUEL SYSTEM.

Provisions are installed for connection to longrange fuel cells.

Section V. FLIGHT CONTROLS

2-37. DESCRIPTION.

The aircraft's primary flight control systemsconsist of conventional rudder, elevator and aileroncontrol surfaces. These surfaces are manually operatedfrom the cockpit through mechanical linkage using acontrol wheel for the ailerons and elevators, andadjustable rudder/brake pedals for the rudder. Both thepilot and copilot have flight controls. Trim control for therudder, elevator and ailerons is accomplished through amanually actuated cable-drum system for each set ofcontrol surfaces. The autopilot has provisions forcontrolling the position of the ailerons, elevators, andrudder. Chapter 3 describes the operation of the autopilotsystem.

2-38. CONTROL WHEELS.

Elevator and aileron control surfaces are operatedby manually actuating either the pilot's or copilot's controlwheel (fig. 2-16). Switches are installed in the grips ofeach control wheel for operation of pitch trim microphone,autopilot disconnect, transponder identification, and chaffdispenser. A manually wound 8-day clock is installed inthe center of the pilot's control wheel, and a digitalclock/timer is installed in the center of the copilot's controlwheel. A map light switch is mounted adjacent to eachclock. For information on operation of the digitalclock/timer, refer to Section XI.

2-39. RUDDER SYSTEM.

a. Rudder Pedals. Aircraft rudder control and nosewheel steering is accomplished by actuation of the rudderpedals from either pilot's or copilot's station (fig. 2-9).The rudder pedals may be individually adjusted in either a

forward or aft position to provide adequate leg room forthe pilot and copilot. Adjustment is accomplished bydepressing the lever alongside the rudder pedal arm andmoving the pedal forward or aft until the locking pinengages in the selected position.

b. Yaw Damp System. A yaw damp system isprovided to aid the pilot in maintaining directional stabilityand increase ride comfort. The system may be used atany altitude and is required for flight above 17,000 feet. Itmust be deactivated for takeoff and landing. The yawdamp system is a part of the autopilot. Operatinginstructions for this system are contained in Chapter 3.The system is controlled by a YAW DAMP switchadjacent to the ELEV TRIM switch on the pedestalextension.

c. Rudder Boost System. A rudder boost system isprovided to aid the pilot in maintaining directional stabilityresulting from an engine failure or a large variation ofpower between the engines. Incorporated in the ruddercable system are two pneumatic rudder boosting servosthat actuates the cables to provide rudder pressure tohelp compensate for asymmetrical thrust.

(1) During operation, a differential pressurevalve accepts bleed air pressure from each engine.When the pressure varies between the bleed air systems,the shuttle in the differential pressure valve moves towardthe low pressure side. As the pressure difference reachesa preset tolerance, a switch closes on the low pressureside which activates the rudder boost system. Thissystem is designed only to help compensate forasymmetrical thrust. Appropriate trimming is to beaccomplished by the pilot. Moving

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Figure 2-16. Control Wheels

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TM 55-1510-219-10

either or both of the bleed air valve switches on theoverhead control panel to PNEU & ENVIRO OFF positionwill disengage the rudder boost system.

NOTECondition levers must be in LOWIDLE position to perform rudderboost check.

(2) The system is controlled by a switchlocated on the extended pedestal placarded RUDDERBOOST ON OFF, and is to be turned on before flight. Apreflight check of the system can be performed during therun-up by retarding the power on one engine to idle andadvancing power on the opposite engine until the powerdifference between the engines is great enough toactivate the switch to turn on the rudder boost system.Movement of the appropriate rudder pedal (left engineidling, right rudder pedal moves forward) will be notedwhen the switch closes, indicating the system isfunctioning properly for low engine power on that side.Repeat the check with opposite power settings to checkfor movement of the opposite rudder pedal. The systemis protected by a 5-ampere circuit breaker placardedRUDDER BOOST, located on the overhead circuitbreaker panel (fig. 2-26).

NOTEWith brake deice on, rudder boostmay be inoperative.

2-40. FLIGHT CONTROLS LOCK.

CAUTIONRemove control locks beforetowing the aircraft or startingengines. Serious damage couldresult in the steering linkage iftowed by a tug with the rudder lockinstalled.

Positive locking of the rudder, elevator andaileron control surfaces, and engine controls (powerlevers, propeller levers, and condition levers) is providedby a removable lock assembly (fig. 2-17) consisting oftwo pins, and an elongated U-shaped strap interconnectedby a chain. Installation of the controls lock isaccomplished by inserting the U-shaped strap around thealigned control levers from the copilot's side; then theaileron/elevator locking pin is inserted through a guidehole in the top of the pilot's control column assembly, thuslocking the control wheel. The rudder is held in a neutralposition by an L-shaped pin which is installed through aguide hole in the floor aft of the pilots rudder pedals. Therudder pedals must be centered to align the hole in the

rudder bellcrank with the guide hole in the floor. Removethe locks in reverse order, i. e. , rudder pin, controlcolumn pin, and power control clamp.

2-41. TRIM TABS.

Trim tabs are provided for all flight controlsurfaces. These tabs are manually activated, and aremechanically controlled by a cable-drum and jackscrewactuator system, except the right aileron tab which is ofthe fixed bendable type. Elevator and aileron trim tabsincorporate neutral, anti-servo action, i. e. , as theelevators or ailerons are displaced from the neutralposition, the trim tab maintains an "as adjusted" position.The rudder trim tab incorporates anti-servo action, i. e. ,as the rudder is displaced from the neutral position thetrim tab moves in the same direction as the controlsurface. This action increases control pressure as rudderis deflected from the neutral position.

a. Elevator Trim Tab Control. The elevator trim tabcontrol wheel placarded ELEVATOR TAB DOWN, UP, ison the left side of the control pedestal and controls a trimtab on each elevator (fig. 210). The amount of elevatortab deflection in degrees from a neutral setting, isindicated by a position arrow.

b. Electric Elevator Trim. The electric elevator trimsystem is controlled by an ELEV TRIM PUSH ON PUSHOFF switch located on the pedestal, dual element thumbswitches on the control wheels, a trim disconnect switchon each control wheel and a circuit breaker on theoverhead circuit breaker panel. The PUSH ON -PUSHOFF switch must be in the ON position to operate thesystem. The dual element thumb switch is movedforward for trimming nose down, aft for nose up, andwhen released returns to the center (off) position. Anyactivation of the trim system through the copilot's trimswitch can be cancelled by activation of the pilot's switch.Operating the pilot's and copilot's switches in opposingdirections simultaneously results in no trim action. Apreflight check of the switches should be accomplishedbefore flight by moving the switches individually on bothcontrol wheels. No one switch alone should operate thesystem; operation of elevator trim should occur only bymovement of pairs of switches. The trim systemdisconnect is a bi-level, push button, momentary typeswitch, located on the outboard grip of each control wheel.Depressing the switch to the first of two levels disconnectsthe autopilot and yaw damp system, and the second leveldisconnects the electric trim

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Figure 2-17. Control Locks

system. The system can be reset by pressing the ONOFF switch on the pedestal to ON again.

c. Aileron Trim Tab Control. The aileron trim tabcontrol, placarded AILERON TAB -LEFT, RIGHT, is onthe control pedestal and will adjust the left aileron trim tabonly (fig. 2-10). The amount of aileron tab deflection,from a neutral setting, as indicted by a position arrow, isrelative only and is not in degrees. Full travel of the tabcontrol moves the trim tab 7-1/2 degrees up and down.

d. Rudder Trim Tab Control. The rudder trim tabcontrol knob, placarded RUDDER TAB LEFT, RIGHT, ison the control pedestal, and controls adjustment of therudder trim tab (fig. 2-10). The amount of rudder tabdeflection, in degrees from a neutral setting, is indicatedby a position arrow.

2-42. WING FLAPS.

The all-metal slot-type wing flaps are electricallyoperated and consist of two sections for each wing.These sections extend from the inboard end of eachaileron to the junction of the wing and fuselage. Duringextension, or retraction, the flaps are operated as a singleunit, each section being actuated by a separate jackscrewactuator. The actuators are driven through flexible shaftsby a single, reversible electric motor. Wing flapmovement is indicated in percent of travel by a flap

position indicator on the forward control pedestal. Fullflap extension and retraction time is approximately 11seconds. The flap control switch is located on the controlpedestal. No emergency wing flap actuation system isprovided. With flaps extended beyond the APPROACHposition, the landing gear warning horn will sound, unlessthe landing gear is down and locked. The circuit isprotected by a 20-ampere circuit breaker, placarded FLAPMOTOR POWER, located on the overhead circuit breakerpanel (fig. 2-26).

a. Wing Flap Control Switch. Flap operation iscontrolled by a three-position switch with a flapshapedhandle on the control pedestal (fig. 2-10). The handle ofthis switch is placarded FLAP and switch positions areplacarded: FLAP UP, APPROACH, and DOWN. Theamount of downward extension of the flaps is establishedby position of the flap switch, and is as follows: UP 0%,APPROACH 40%, and DOWN 100%. Limit switches,mounted on the right inboard flap, control flap travel. Theflap control switch, limit switch, and relay circuits areprotected by a 5-ampere circuit breaker, placarded FLAPCONTR located on the overhead circuit breaker panel(fig. 2-26). Flap positions between UP and APPROACHcannot be selected. For intermediate flap positionsbetween AP-

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TM 55-1510-219-10

PROACH and DOWN, the APPROACH position acts asan off position. To return the flaps to any positionbetween full DOWN and APPROACH, place the flapswitch to UP and when desired flap position is obtained,return the switch to the APPROACH detent. In theevent that any two adjacent flap sections extend 3 to 5degrees out of phase with the other, a safetymechanism is provided to discontinue power to the flapmotor.

b. Wing Flap Position Indicator. Flapposition in percent of travel from "O" percent (UP) to100 percent (DOWN), is shown on an indicator,placarded FLAPS, located on the subpanel (fig. 2-6).The approach and full down or extended flap position is14 and 34 degrees, respectively. The flap positionindicator is protected by a 5-ampere circuit breaker,placarded FLAP CONTR, located on the overheadcircuit breaker panel (fig. 2-26).

Section VI. PROPELLERS

2-43. DESCRIPTION.

A three blade aluminum propeller is installed oneach engine. The propeller is of the full feathering,constant speed, counter-weighted, reversible type,controlled by engine oil pressure through single action,engine driven propeller governors. The propeller is flangemounted to the engine shaft. Centrifugal counterweights,assisted by a feathering spring, move the blades towardthe low RPM (high pitch) position and into the featheredposition. Governor boosted engine oil pressure movesthe propeller to the high RPM (low pitch) hydraulic stopand reverse position. The propellers have no low RPM(high pitch) stops; this allows the blades to feather afterengine shutdown. Low pitch propeller position isdetermined by the low pitch stop which is a mechanicallyactuated, hydraulic stop. Beta and reverse blade anglesare controlled by the power levers in the beta and reverserange.

2-44. FEATHERING PROVISIONS.

Both manual and automatic propeller featheringsystems are provided. Manual feathering is accomplishedby pulling the corresponding propeller lever aft past afriction detent. To unfeather, the propeller lever is pushedforward into the governing range. An automaticfeathering system, will sense loss of torque and willfeather an unpowered propeller. Feathering springs willfeather the propeller when it is not turning.

a. Automatic Feathering. The automatic featheringsystem provides a means of immediately dumping oilfrom the propeller servo to enable the feathering springand counterweights to start feathering action of the bladesin the event of an engine failure. Although the system isarmed by a switch on the overhead control panel,placarded AUTOFEATHER ARM OFF TEST, thecompletion of the arming phase occurs when both powerlevers are advanced above 90% Ni at which time bothindicator lightson the caution/advisory annunciator panelindicate a fully armed system. The annunciator lights are

green and are placarded #1 AUTOFEATHER (left engine)and #2 AUTOFEATHER (right engine). The system willremain inoperative as long as either power lever isretarded below 90% Ni position, unless TEST position ofthe AUTOFEATHER SWITCH is selected to disable thepower lever limit switches. The system is designed foruse only during takeoff and landing and should be turnedoff when establishing cruise climb. During takeoff orlanding, should the torque for either engine drop to anindication between 16 21%, the autofeather system forthe opposite engine will be disarmed. Disarming isconfirmed when the AUTOFEATHER light of the oppositeengine becomes extinguished. If torque drops further, toa reading between 9 and 14%, oil is dumped from theservo of the effected propeller allowing a featheringspring and counter-weights to move the blades intofeathered position. Feathering also causes theAUTOFEATHER light of the feathered propeller toextinguish. At this time, both annunciatorAUTOFEATHER lights are extinguished, the propeller ofthe defective engine has feathered, and the propeller ofthe operative engine has been disarmed from theautofeathering capability. Only manual feathering controlremains for the second propeller.

b. Propeller Autofeather Switch. Autofeathering iscontrolled by an AUTOFEATHER switch on the overheadcontrol panel (fig. 2-18). The threeposition switch isplacarded ARM, OFF and TEST, and is spring-loadedfrom TEST to OFF. The ARM position is used only duringtakeoff and landing. The TEST position of the switch,enables the pilot to check readiness of the autofeathersystems, below 88% to 92% N1, and is for groundcheckout purposes only.

c. Autofeather Lights. Two green lights on thecaution/advisory annunciator panel, placardedAUTOFEATHER #1 and #2. When illuminated, the lightsindicate that the autofeather system is armed. Both lightswill be extinguished if either propeller has beenautofeathered or if the system is disarmed by retarding apower lever. Autofeather circuits are protected by one 5-ampere circuit breaker placarded

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AUTO FEATHER, located on the overhead circuit breakerpanel (fig. 2-26).

2-45. PROPELLER GOVERNORS.

Two governors (a constant speed governor, andan overspeed governor) control propeller RPM. Theconstant speed governor, mounted on top of the propellerreduction housing, controls the propeller through its entirerange. The propeller control lever operates the propellerby means of this governor. If the constant speedgovernor should malfunction and request more than 2000RPM, the overspeed governor cuts in at 2080 RPM anddumps oil from the propeller to limit RPM. A solenoid,actuated by the PROP GOV TEST switch located on theoverhead control panel (fig. 2-18), is provided to resetthe overspeed governor to approximately 1830 to 1910RPM for test purposes. If the propeller sticks or movestoo slowly to prevent an overspeed condition, the powerturbine governor, contained within the constant speedgovernor housing, acts as a fuel topping governor. Whenthe propeller reaches 2120 RPM, the fuel toppinggovernor limits the fuel flow to the gas generator, therebyreducing the power driving the propeller. Duringoperation in the reverse range, the power turbinegovernor is reset to approximately 95% propeller RPMbefore the propeller reaches a negative pitch angle. Thisinsures that the engine power is limited to maintain apropeller RPM of somewhat less than that of the constantspeed governor setting. The constant speed governorwill, therefore, always sense an underspeed condition anddirect oil pressure to the propeller servo piston to permitpropeller operation in beta and reverse range.

2-46. PROPELLER TEST SWITCHES.

Two two-position switches on the overheadcontrol panel (fig. 2-18), are provided for operationaltesting of the propeller systems. Placarding above theswitches reads PROP GOV TEST. Each switch controlstest circuits for the corresponding propeller. In the testposition, the switches are used to test the function of thecorresponding overspeed governor. Refer to Chapter 8,for test procedure. Propeller test circuits are protected byone 5-ampere circuit breaker placarded PROP GOV.located on the overhead circuit breaker panel (fig. 2-26).

2-47. PROPELLER SYNCHROPHASER.

a. Operation. The propeller synchrophaserautomatically matches the RPM of the right propeller(slave propeller) to that of the left propeller (masterpropeller) and maintains the blades of one propeller at apredetermined relative position with the blades of theother propeller. To prevent the right propeller from losing

excessive RPM if the left propeller is feathered while thesynchrophaser is on, the synchrophaser has a limitedrange of control from the manual governor setting.Normal governor operation is unchanged but thesynchrophaser will continuously monitor propeller RPMand reset the governor as required. A magnetic pickupmounted in each propeller overspeed governor andadjacent to each propeller deice brush block transmitselectric pulses to a transistorized control box installedforward of the pedestal. The right propeller RPM andphase will automatically be adjusted to correspond to theleft. To change RPM, adjust both propeller controls at thesame time. This will keep the right governor setting withinthe limiting range of the left propeller. If thesynchrophaser is on but is unable to adjust to the rightpropeller to match the left, the actuator has reached theend of its travel. To recenter, turn the switch off,synchronize the propellers manually, and turn the switchback on.

b. Control Box. The control box converts any pulserate differences into correction commands, which aretransmitted to a stepping type actuator motor mounted onthe right engine cowl forward support ring. The motorthen trims the right propeller governor through a flexibleshaft and trimmer assembly to exactly match the leftpropeller. The trimmer, installed between the governorcontrol arm and the control cable, screws in or out toadjust the governor while leaving the control lever settingconstant. A toggle switch installed adjacent to thesynchrophaser turns the system on. With the switch off,the actuator automatically runs to the center of its rangeof travel before stopping to assure normal function whenused again. To operate the system, synchronize thepropeller in the normal manner and turn thesynchrophaser on. The system is designed for in-flightoperations and is placarded to be off for take-off andlanding. Therefore, with the system on and the landinggear extended, the master caution lights will illuminateand a yellow light on the caution/advisory annunciatorpanel, PROP SYNC ON, will illuminate.

c. Synchroscope. The propeller synchroscope,provides an indication of synchronization of thepropellers. If the right propeller is operating at a higherRPM than the left, the face of the synchroscope, a blackand white cross pattern, spins in a clockwise rotation.Left, or counterclockwise, rotation indicates a higher RPMof the left propeller. This instrument aids the pilot inobtaining complete synchronization of propellers. Thesystem is protected by a 5-ampere circuit breakerplacarded PROP SYNC, located on the overhead circuitbreaker panel (fig. 2-26).

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2-48. PROPELLER LEVERS.

Two propeller levers on the control pedestal (fig,2-10), placarded PROP, are used to regulate propellerspeeds. Each lever controls a primary governor, whichacts to regulate propeller speeds within the normaloperation range. The full forward position of the levers isplacarded TAKEOFF, LANDING AND REVERSE andalso HIGH RPM. The full aft position of the levers isplacarded FEATHER. When a lever is placed at HIGHRPM, the propeller may attain a static RPM of 2,000depending upon power lever position. As a lever ismoved aft, passing through the propeller governing range,but stopping at the feathering detent, propeller RPM willcorrespondingly decrease to the lowest limit. Moving apropeller lever aft past the detent into FEATHER willfeather the propeller.

2-49. PROPELLER REVERSING.

CAUTIONDo not move the power levers intoreverse range without the enginerunning. Damage to the reverselinkage mechanisms will occur.

CAUTIONPropeller reversing on unimprovedsurfaces should be accomplishedcarefully to prevent propellererosion from reversed airflow and,

in dusty conditions, to preventobscuring the operator's vision.

CAUTIONTo prevent an asymmetrical thrustcondition, propeller levers must bein HIGH RPM position prior topropeller reversing.

The propeller blade angle may be reversed toshorten landing roll. To reverse, propeller levers must bepositioned at HIGH RPM (full forward), and the powerlevers are lifted up to pass over the IDLE detent, thenpulled aft into REVERSE setting. One yellow cautionlight, placarded REV NOT READY, on thecaution/advisory annunciator panel, alerts the pilot not toreverse the propellers. This light illuminates only whenthe landing gear handle is down, and if propeller leversare not at HIGH RPM (full forward). This circuit isprotected by a 5ampere circuit breaker, placardedLANDING GEAR RELAY, located on the overhead circuitbreaker panel (fig. 2-26).

2-50. PROPELLER TACHOMETERS.

Two tachometers on the instrument panel registerpropeller speed in hundreds of RPM (fig. 2-28). Eachindicator is slaved to a tachometer generator unitattached to the corresponding engine.

Section VII. UTILITY SYSTEMS

2-51. DEFROSTING/DEFOGGING SYSTEM.

a. Description. The defrosting/defogging system isan integral part of the heating and ventilation system.The system consists of two warm air outlets connected byducts to the heating system. One outlet is just below thepilot's windshield and the other is below the copilot'swindshield. A push-pull control, placarded DEFROSTAIR, on the pilot's subpanel, manually controls airflow tothe windshield. When pulled out, defrosting air is ductedto the windshield. As the control is pushed in, there is acorresponding decrease in airflow.

b. Automatic Operation.

1. Vent blower switches As required.

2. Cabin temperature mode selector switchAUTO

3. Cabin temperature control As required.

4. Cabin air, copilot air, pilot air, and defrostair controls As required.

c. Maximum Windshield Defrosting.

1. Pilot air, copilot air IN.

2. Cabin air and defrost air controls Out

3. Cabin temperature mode selectorswitch MAN HEAT.

4. Cold air outlets As required.

5. Manual temperature switch As required.

d. Manual Operation. If the automatic temperaturecontrol should fail to operate, the tempera-

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ture (of defrost air and cabin air) may be controlledmanually by manipulating the CABIN TEMP MODEcontrol switch between the OFF and MAN HEATpositions.

2-52. SURFACE DEICER SYSTEM.

a. Description. Ice accumulation is removed fromeach inboard and outboard wing leading edge, and bothhorizontal stabilizers by the flexing of deicer boots whichare pneumatically actuated. Engine bleed air, from theengine compressor, is used to supply air pressure toinflate the deicer boots, and to supply vacuum, throughthe ejector system, for boot hold down during flight. Apressure regulator protects the system from over inflation.When the system is not in operation, a distributor valveapplies vacuum to the boots for hold down.

CAUTIONOperation of the surface deicesystem in ambient temperaturesbelow -40°C can cause permanentdamage to the deice boots.

b. Operation.

(1) Deice boots are intended to remove iceafter it has formed rather than prevent its formation. Forthe most effective deicing operation, allow at least 1/2inch of ice to form on the boots before attempting iceremoval. Very thin ice may crack and cling to the bootsinstead of shedding.

NOTENever cycle the system rapidly, thismay cause the ice to accumulateoutside the contour of the inflatedboots and prevent ice removal.

(2) A three position switch on the overheadcontrol panel placarded DEICE MANUAL OFF SINGLECYCLE AUTO, controls the deicing operation. The switchis spring loaded to return to the OFF position fromSINGLE CYCLE AUTO or MANUAL. When the SINGLECYCLE AUTO position is selected, the distributor valveopens to inflate the wing boots. After an inflation periodof approximately 6 seconds, an electronic timer switchesthe distributor to deflate the wing boots and a 4 secondinflation begins in the horizontal stabilizer boots. Whenthese boots have inflated and deflated, the cycle iscomplete.

(3) If the switch is held in the MANUALposition, the boots will inflate simultaneously and remaininflated until the switch is released. The switch will return

to the OFF position when released. After the cycle, theboots will remain in the vacuum hold down condition untilagain actuated by the switch.

(4) Either engine is capable of providingsufficient bleed air for all requirements of the surfacedeicer system. Check valves in the bleed air and vacuumlines prevent backflow through the system during single-engine operation. Regulated pressure is indicated on agage, placarded PNEUMATIC PRESSURE, located onthe copilots subpanel.

2-53. ANTENNA DEICE SYSTEM.

a. Description. The antenna deice system removesice accumulation from the inboard wing dipole antennasand the aft rotating boom dipole antenna. Pressureregulated bleed air from the engines supplies pressure toinflate the boots. To assure operation of the system in theevent of failure of one engine, a check valve isincorporated in the bleed air line from each engine toprevent loss of pressure through the compressor of theinoperative engine. Inflation and deflation phases arecontrolled by a distributor valve. Deice boots areintended to remove ice after is has formed rather than toprevent it's formation. For the most effective deicingoperation, allow at least 1/8 to 1/4 inch of ice to form onthe boots before attempting ice removal. Very thin icemay crack and cling to the boots instead of shedding.

NOTENever cycle the system rapidly.This may cause the ice toaccumulate outside the contour ofthe inflated boots and prevent iceremoval.

b. Antenna Deice System Switch. The antennadeice system is controlled by a switch placarded ANTDEICE, SINGLE-OFF-MANUAL located on the overheadcontrol panel (fig. 2-18). The switch is spring loaded toreturn to the OFF position from the SINGLE or MANUALposition. When the switch is set to the single position, thesystem will run through one timed inflation-deflation cycle.When the switch is held in the MANUAL position theboots will inflate and remain inflated until the switch isreleased.

c. Forward Wide Band Data Link Antenna RadomeAnti-Ice. The forward wide band data link antennaradome anti-ice system utilizes engine bleed air toprevent the formation of ice on the radome.

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The system is controlled by a switch placarded RADOMEANTI-ICE-ON located on the overhead control panel. Atemperature sensing element installed in the dischargeduct measures the hot air temperature as it leaves themixing chamber. When the hot air temperature reaches130 degrees Fahrenheit, the green RADOME HEAT lightin the mission caution/ advisory panel will illuminate. Ifthe air temperature exceeds 200 degrees Fahrenheit, aheat detection switch, located adjacent to the temperaturesensing element, will close the solenoid shutoff valves,and the yellow RADOME HOT light located on themission caution/advisory panel will illuminate. Theshutoff valves will automatically open after the airtemperature cools to approximately 130 degreesFahrenheit. The system is protected by a 7. 5 amperecircuit breaker placarded RADOME ANTI-ICE located onthe overhead control panel.

2-54. PROPELLER ELECTROTHERMAL ANTI-ICESYSTEM.

a. Description. Electrothermal anti-ice boots arecemented to each propeller blade to prevent ice formationor to remove the ice from the propellers. Each thermalboot consists of one outboard and one inboard heatingelement, and receives electrical power from the deicetimer. This timer sends current to all propeller deiceboots and prevents the boots from overheating by limitingthe time each element is energized. Four intervals ofapproximately 30 seconds each complete one cycle.Current consumption is monitored by a propeller ammeteron the copilot's subpanel. Two 20-ampere circuitbreakers placarded PROP ANTI ICE LEFT and RIGHTand 5-ampere propeller control circuit breaker placardedPROP ANTI-ICE CONTR on the overhead circuit breakerpanel (fig. 2-26), protect the propeller electrothermaldeice system during manual operation. A 25 amperecircuit breaker placarded PROP ANTI-ICE AUTO,protects the system in automatic operation.

b. Automatic Operation. A control switch on theoverhead control panel placarded PROP OFF AUTO isprovided to activate the automatic system. A deiceammeter on the center subpanel registers the amount ofcurrent (14 to 18 amperes) passing through the systembeing used. During AUTO operation, power to the timerwill be cut off if the current rises above 25 amperes.Current flows from the timer to the brush assembly andthen to the slip rings installed on the spinner backingplate. The slip rings carry the current to the deice bootson the propeller blades. Heat from the boots reduces thegrip of the ice which is then thrown off by centrifugalforce, aided by the air blast over the propeller surfaces.

Power to the two heating elements on each blade, theinner and outer element, is cycled by the timer in thefollowing sequence: right propeller outer element, rightpropeller inner element, left propeller outer element, leftpropeller inner element. Loss of one heating elementcircuit on one side does not mean that the entire systemmust be turned off. Proper operation can be checked bynoting the correct level of current usage on the ammeter.An intermittent flicker of the needle approximately each30 seconds indicates switching to the next group ofheating elements by the timer.

c. Manual Operation. The manual propeller deicesystem is provided as a backup to the automatic system.A control switch located on the overhead control panel,placarded PROP INNER OUTER, controls the manualoverride relays. When the switch is in the OUTERposition, the automatic timer is overridden and power issupplied to the outer heating elements of both propellerssimultaneously. The switch is of the momentary type andmust be held in position until the ice has been dislodgedfrom the propeller surface. After deicing with the outerelements, the switch is to be held in the INNER position toperform the same function for the inner elements of bothpropellers. The loadmeters will indicate approximately a5% increase of load per meter when manual prop deice isoperating. The prop deice ammeter will not indicate anyload in the manual mode of operation.

2-55. PITOT AND STALL WARNING HEAT SYSTEM.

CAUTIONPitot heat should not be used formore than 15 minutes while theaircraft is on the ground.Overheating may damage theheating elements.

a. Pitot Heat. Heating elements are installed in thepitot masts located on the nose. Each heating element iscontrolled by an individual switch placarded PITOT ONLEFT or RIGHT, located on the overhead control panel(fig. 2-18). It is not advisable to operate the pitot heatsystem on the ground except for testing or for shortintervals of time to remove ice or snow from the mast.Circuit protection is provided by two 7. 5 amperes circuitbreakers, placarded PITOT HEAT, on the overhead circuitbreaker panel (fig. 2-26).

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CAUTIONThe heating elements protect thestall warning lift transducer vaneand face plate from ice, however, abuildup of ice on the wing maychange or disrupt the airflow andprevent the system from accuratelyindicating an imminent stall.

b. Stall Warning Heat. The lift transducer isequipped with anti-icing capability on both the mountingplate and the vane. The heat is controlled by a switchlocated on the overhead control panel placarded STALLWARN. The level of heat is minimal for ground operationbut is automatically increased for flight operation throughthe left landing gear saety switch. Circuit protection isprovided by a 15-ampere circuit breaker, placardedSTALL WARN, on the overhead circuit breaker panel (fig.2-26).

2-56. PITOT AND STATIC SYSTEM.

a. Description. The pitot and static system suppliesstatic pressure to two airspeed indicators, two altimeters,two vertical velocity indicators, and ram air to theairspeed indicators. This system consists of two pitotmasts (one located on each side of the lower position ofthe nose), static air pressure ports in the aircraft's exteriorskin on each side of the aft fuselage, and associatedsystem plumbing. The pitot head is protected from iceformation by internal electric heating elements.

b. Alternate Static Air Source. An alternate static airline, which terminates just aft of the rear pressurebulkhead, provides a source of static air for the pilot'sinstruments in the event of source failure from the pilot'sstatic air line. A control on the pilot's subpanel placardedPILOTS STATIC AIR SOURCE, may be actuated toselect either the NORMAL or ALTERNATE air source bya two position selector valve. The valve is secured in theNORMAL position by a spring clip. Refer to Chapter 7 forairspeed indicator and altimeter calibration informationwhen using the alternate air source.

c. Stall Warning System. The stall warning systemconsists of a transducer, a lift computer, a warning horn,and a test switch. Angle of attack is sensed byaerodynamic pressure on the lift transducer vane locatedon the left wing leading edge. When a stall is imminent,the output of the transducer activates a stall warning horn.The system has preflight test capability through the use ofa switch placarded STALL WARN TEST OFFLDG GEARWARN TEST on the right subpanel. Holding this switch inthe STALL WARN TEST position actuates the warninghorn by moving the transducer vane. The circuit is

protected by a 5-ampere circuit breaker, placarded STALLWARN, on the overhead circuit breaker panel.

2-57. BRAKE DEICE SYSTEM.

a. Description. A heated-air brake deice systemmay be used on the ground or in flight with gear retractedor extended. When activated, hot air is diffused bymeans of a manifold assembly over the brake discs ineach wheel. Manual and automatic controls are provided.There are two primary occasions which require brakedeicing. The first is when an aircraft has been parked in afreezing atmosphere allowing the brake systems tobecome contaminated by freezing rain, snow, or ice, andthe aircraft must be moved or taxied. The secondoccasion is during flight through icing conditions with wetbrake assemblies presumed to be frozen, which must bethawed prior to landing to avoid possible tire damage andloss of directional control. Hot air for the brake deicesystem comes from the compressor stage of both enginesobtained by means of a solenoid valve attached to thebleed air system which serves both the surface deicesystem and the pneumatic systems operation.

b. Operation. A switch on the overhead controlpanel, placarded BRAKE DEICE, controls the solenoidvalve by routing power through a control module boxunder the aisleway floorboards. When the switch is on,power from a 5-ampere circuit breaker on the overheadcircuit breaker panel is applied to the control module. A10-minute timer limits operation and avoids excessivewheel well temperatures when the landing gear isretracted. The control module also contains a circuit tothe green BRAKE DEICE ON annunciator light, and has aresetting circuit interlocked with the gear uplock switch.When the system is activated, the BRAKE DEICE ONlight should be monitored and the control switch selectedOFF after the light extinguishes otherwise, on the nextgear extension the system will restart without pilot action.The control switch should also be selected OFF, if deiceoperation fails to self-terminate after 10 minutes. If theautomatic timer has terminated brake deicer operationafter the last retraction of the landing gear, the landinggear must be extended in order to obtain further operationof the system.

(1) The BLEED AIR FAIL lights maymomentarily illuminate during simultaneous operation ofthe surface deice and brake deice systems at low N1speeds. If the lights immediately extinguish. this may bedisregarded.

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(2) During certain ambient conditions, use ofthe brake deice system may reduce available enginepower, and during flight will result in a TGT rise ofapproximately 20°C. Appropriate performance chartsshould be consulted before brake deice system use. Ifspecified power cannot be obtained without exceedinglimits, the brake deice system must be selected off untilafter takeoff is completed. TGT limitations must also beobserved when setting climb and cruise power. Thebrake deice system is not to be operated above 15°Cambient temperature except to test the system. Thesystem is not to be operated for longer than 10 minutes(one deicer cycle) with the landing gear retracted. Ifoperation does not automatically terminate afterapproximately 10 minutes following gear retraction, thesystem must be manually selected off. During periods ofsimultaneous brake deice and surface deice operation,maintain 85% Nl or higher. If inadequate pneumaticpressure is developed for proper surface deicer bootinflation, select the brake deice system off. Both sourcesof pneumatic bleed air must be in operation during brakedeice system use. Select the brake deice system offduring single-engine operation. Circuit protection isprovided by a 5-ampere circuit breaker, placardedBRAKE DEICE, on the overhead circuit breaker panel(fig. 2-26).

2-58. FUEL SYSTEM ANTI-ICING.a. Description. An oil-to-fuel heat exchanger,

located on each engine accessory case, operatescontinuously and automatically to heat the fuel sufficientlyto prevent freezing of any water in the fuel. No controlsare involved. Two external fuel vents are provided oneach wing. One is recessed to prevent ice formation; theother is electrically heated and is controlled by two toggleswitches on the overhead control panel placarded FUELVENT ON, LEFT or RIGHT (fig. 2-18). They areprotected by two 5-ampere circuit breakers, placardedFUEL VENT HEAT, RIGHT or LEFT, located on theoverhead circuit breaker panel (fig. 2-26). Each fuelcontrol unit's pneumatic line is protected against ice by anelectrically heated jacket protected by a 7. 5amperecircuit breaker located on the overhead circuit breakerpanel placarded FUEL CONTR HEAT, LEFT or RIGHT(fig. 2-26).

CAUTIONTo prevent overheat damage toelectrically heated anti-ice jackets,the FUEL VENT heat switchesshould not be turned ON unlesscooling air will soon pass over thejackets.

b. Normal Operation. For normal operation,switches for the FUEL VENTS anti-ice circuits are turnedON as required during the BEFORE TAKEOFFprocedures (Chapter 8).

2-59. WINDSHIELD ELECTROTHERMAL ANTIICESYSTEM.

a. Description. Both pilot and copilot windshieldsare provided with an electrothermal anti-ice system. Eachwindshield is part of an independent electrothermal anti-ice system. Each system is comprised of the windshieldassembly with heating wires sandwiched between glasspanels, a temperature sensor attached to the glass, anelectrothermal controller, two relays, a control switch, andtwo circuit breakers. Two switches, placarded WSHLDANTIICE NORMAL -OFF HI PILOT, COPILOT, locatedon the overhead control panel (fig. 2-18) control systemoperation. Each switch controls one electrothermalwindshield system. The circuits of each system areprotected by a 5-ampere circuit breaker and a 50-amperecircuit breaker which are not accessible to the flight crew.The 50-ampere circuit breakers are located in the powerdistribution panel under the floor ahead of the main spar.The 5-ampere circuit breakers are located on panelsforward of the instrument panel.

b. Normal Operation. Two levels of heat areprovided through the three position switches placardedNORMAL in the aft position, OFF in the center position,and HI after lifting the switch over a detent and moving itto the forward position. In the NORMAL position, heat isprovided for the major portion of each windshield. In theHI position, heat is provided at a higher watt density to asmaller portion of the windshield. The lever lock featureprevents inadvertent switching to the HI position duringsystem shutdown.

2-60. PRESSURIZATION SYSTEM.a. Pressure Differential. The pressure vessel is

designed for a normal working pressure differential of 6.0 PSI, which will provide a cabin pressure altitude of 3870feet at an aircraft altitude of 20,000 feet, and a nominalcabin altitude of 9840 feet at an aircraft altitude of 31,000feet.

b. Description. A mixture of bleed air from theengines, and ambient air, is available for pressurization tothe cabin at a rate of approximately 10 to 17 pounds perminute. Approximately 85% NI is required whenoperating with one engine. The flow control unit of eachengine controls the bleed air from the engine to make itusable for pressurization by mixing ambient air with thebleed air depending

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TM 55-1510-219-10

upon aircraft altitude and ambient temperature. Ontakeoff, excessive pressure bumps are prevented bylanding gear safety switch actuated solenoidsincorporated in the flow control units. These solenoids,through a time delay, stage the input of ambient air flowby allowing ambient air flow introduction through the leftflow control unit first, ten seconds later, air flow throughthe right flow control unit. The Bleed Air Switches,located on the overhead control panel (fig. 2-1 8) operatean integral electric solenoid which controls the bleed air tothe firewall shutoff valves.

c. Cabin Altitude and Rate-of-Climb Controller. Acontrol panel is installed on the copilot's inboard subpanelfor operation of the system. A knob (fig. 2-6), placardedINC RATE controls the rate of change of pressurization.A control, placarded CABIN CONTROLLER is used to setthe desired cabin altitude. For proper cabinpressurization, the CABIN CONTROLLER should be set500 feet above cruise altitude. For landing select 500feet above field pressure altitude. The selected altitude isdisplayed on a mechanically coupled dial above thecontrol, placarded CABIN ALT-FT. Mechanically coupledto the cabin altitude dial, placarded ACFTX1000. Thisdial indicates the maximum altitude the aircraft may beflown at to maintain the desired cabin altitude withoutexceeding the design pressure differential. A switch,placarded CABIN PRESS DUMP-PRESS-TEST, isprovided to control pressurization. The switch is springloaded to the PRESS position. In the DUMP position, the

safety valve will be opened and the cabin will bedepressurized to the aircraft altitude. In the PRESSposition, cabin altitude is controlled by the CABINCONTROLLER control. In the TEST position, the landinggear safety switch is bypassed to enable testing of thepressurization system on the ground. Operatinginstructions are contained in Chapter 8.

d. Cabin Rate-of-Climb Indicator. An indicator,placarded CABIN CLIMB, is installed on the copilot's sideof the instrument panel (fig. 2-28). The cabin rate-of-climb controller is calibrated in thousands-of-feet per-minute change in cabin altitude.

e. Cabin Altitude Indicator. An indicator, placardedCABIN ALT, is installed in the instrument panel (fig. 2-28)above the cabin rate-of-climb indicator. The longerneedle indicates aircraft altitude in thousands-of-feet onthe outside dial. The shorter needle indicates pressuredifferential in PSI on the inner dial. Maximum differentialis 6. 1 PSI.

f Outflow Valve. A pneumatically operated outflowvalve, located on the aft pressure bulkhead, maintains theselected cabin altitude and rate-of-climb commanded bythe cabin rate-of-climb and altitude controller on thecopilot's instrument panel.

Figure 2-78. Overhead Control Panel

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TM 55-1510-219-10

As the aircraft climbs, the controller modulates the outflowvalve to maintain a selected cabin rate of climb andincreases the cabin differential pressure until themaximum cabin pressure differential is reached. At acabin altitude of 12,500 feet a pressure switch mountedon the back of the overhead control panel completes acircuit to illuminate a red warning annunciator light, ALTWARN, to warn of operation requiring oxygen. This lightis protected by a 5ampere breaker, placarded PRESSCONTR.

g. Pressurization Safety Valve. Before takeoff, thesafety valve is open with equal pressure between thecabin and the outside air. The safety valve closes onliftoff if the CABIN PRESS CONTR switch on the copilotssubpanel is in the PRESS mode. The safety valveadjacent to the outflow valve provides pressure relief inthe event of failure of the outflow valve. This valve isalso used as a dump valve and is opened by vacuumwhich is controlled by a solenoid valve operated by thecabin pressure dump switch adjacent to the controller. Itis also wired through the right landing gear safety switch.If either of these switches is open, or the vacuum sourceor electrical power is lost, the safety valve will close toatmosphere except at maximum differential pressure of 6.1 PSI. A negative pressure relief diaphragm is alsoincorporated into the outflow and safety valves to preventoutside atmospheric pressure from exceeding cabinpressure during rapid descent.

h. Drain. A drain in the outflow valve static controlline is provided for removal of accumulated moisture.The drain is located behind the lower sidewall upholsteryaccess panel in the baggage section of the aftcompartment.

i. Flow Control Unit. A flow control unit forward ofthe firewall in each nacelle controls bleed air flow and themixing of ambient air to make up the total air flow to thecabin for pressurization, heating, and ventilation. Thebleed air switches located on the overhead control panel(fig. 2-18) operates an integral electric solenoid whichcontrols the bleed air to the firewall shutoff valves. Anormally open solenoid operated by the landing gearsafety switch controls the introduction of ambient air flowto the cabin on takeoff.

(1) The unit receives bleed air from the engineinto an ejector which draws ambient air into the nozzle ofthe venturi. The mixed air is then forced into the bleed airline routed to the cabin.

(2) Bleed air flow is controlled automatically.When the aircraft is on the ground, circuitry from thelanding gear safety switch prevents ambient air fromentering the flow control unit to provide maximumheating.

(3) The bleed air firewall shutoff valve in thecontrol unit is a spring loaded, bellows operated valve thatis held in the open position by bleed air pressure. Whenthe electric solenoid_ is shut off, or when bleed airdiminishes on engine shutdown (in both cases thepressure to the firewall shutoff valve is cut off), thefirewall valve closes.

2-61. OXYGEN SYSTEM.a. Description. The oxygen system (fig. 2-19) is

provided primarily as an emergency system, however, thesystem may be used to provide supplemental (first aid)oxygen. Two 64 cubic foot capacity oxygen supplycylinders charged with aviator's breathing oxygen areinstalled in the unpressurized portion of the aircraft behindthe aft pressure bulkhead. The pilot and copilot positionsare equipped with diluter demand type regulators, whichmix the proper amount of oxygen for a given amount ofair at altitude. Also a first aid oxygen mask is provided inthe cabin. Oxygen system pressure is shown by twogages placarded OXYGEN SUPPLY PRESSURE, locatedaft of the pilot's oxygen regulator control panel. Twopressure reducers, located in the unpressurized portion ofthe aircraft behind the aft bulkhead, lower the pressure inthe system to 400 PSI, and route oxygen to the regulatorcontrol panels. Both cylinders are interconnected, sorefilling can be accomplished through a single filler valvelocated on the aft right side of the fuselage exterior. Apressure gage is mounted in conjunction with the fillervalve, and each cylinder has a pressure gage. Table 2-4shows oxygen duration capacities of the system.

b. Regulator Control Panels. Each regulator controlpanel contains a blinker-type flow indicator, a 500 PSIpressure gage, a red emergency pressure control leverplacarded EMERGENCY NORMAL TEST MASK, a whitediluter control lever placarded 100% OXYGEN -NORMALOXYGEN, and a green supply control lever placarded ONOFF.

(1) Diluter control lever. The diluter controllever selects either normal or 100% oxygen, but acts toselect only when the emergency pressure control lever isin the NORMAL position.

CAUTIONWhen not in use, the diluter controllever should be left in the 100%OXYGEN position to preventregulator contamination.

(2) Emergency pressure control lever. Theemergency pressure control lever has three positions.

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TM 55-1510-219-10

Figure 2-19. Oxygen System Schematic

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TM 55-1510-219-10

Two positions control oxygen consumption for theindividual using oxygen, and the remaining positionserves for testing hose and mask integrity. In theEMERGENCY position, the control lever causes 100%oxygen to be delivered at a safe, positive pressure. In theNORMAL position, the lever allows delivery of normal or100% oxygen, depending upon the selection of the diluter

control lever. In TEST MASK position, 100% oxygen atpositive pressure is delivered to check hose and maskintegrity.

(3) Supply control lever. The supply controllever (green), placarded ON OFF, turns the oxygensupply on or off at the regulator control panel.

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Table 2-4. Oxygen Flow Planning Rates vs Altitude(All Flows in LPM Per Mask at NTPD)

TM 55-1510-219-10

(4) Oxygen supply pressure gage.

WARNINGThe 500 PSI oxygen pressure gageprovided on the oxygen controlpanels should never indicate over400 PSI. If the pressure exceeds400 PSI, a malfunction of thepressure reducer is indicated.

Whenever oxygen is inhaled, a blinker-vaneslides into view within the flow indicator window, showingthat oxygen is being released. When oxygen is exhaled,the blinker vane vanishes from view.

NOTECheck to insure that the OXYGENSUPPLY PRESSURE gage registersadequate pressure before eachflight. When oxygen is in use, acheck of the supply pressureshould be made at intervals duringflight to note the quantity availableand to approximate the supplyduration. The outside temperatureis reduced as an aircraft ascends tohigher altitudes. Oxygen cylindersthus cooled by temperature changewill show a pressure drop. Thistype of drop in pressure will raiseagain upon return to a lower orwarmer altitude. A valid cause foralarm would be the rapid loss ofoxygen pressure when the aircraftis in level flight or descending;should this condition arise,descend as rapidly as possible toaltitude which does not require theuse of oxygen.

WARNINGPure oxygen will supportcombustion. Do not smoke whileoxygen is in use.

WARNINGIf any symptoms occur suggestiveof the onset of hypoxia,immediately set the emergencypressure control lever to theEMERGENCY position and descendbelow 10,000 feet. Whenevercarbon monoxide or other noxiousgas is present or suspected, set thedilutor control lever to 100%

OXYGEN and continue breathingundiluted oxygen until the dangeris past.

c. Oxygen Masks. Oxygen masks for the pilot andcopilot are provided as personal equipment. To connect amask into the oxygen system, the individual connects theline attached to the mask to the flexible hose which isattached to the cockpit sidewall. The microphone in theoxygen mask is provided with a cord for connecting withthe helmet microphone jack. To test mask and hoseintegrity, the individual places the supply control lever onthe regulator control panel to the ON position, puts on andadjusts his mask, selects TEST MASK position, andchecks for leaks.

d. Normal Operation. Oxygen pressure ismaintained at all times to the regulator control panels ifthe cylinder shut-off valves are on and if there is pressurein the cylinders. Each individual places the supply lever(green) on his regulator control panel to the ON position,and the diluter lever (white) to the NORMAL OXYGENposition.

e. Emergency Operation. For emergency operation,the affected crew member selects the EMERGENCYposition of the emergency pressure control lever (red) onhis regulator control panel. This selection provides 100%oxygen at a positive pressure, regardless of the positionof the diluter control lever on his panel.

f First Aid Operation. A first aid oxygen mask isinstalled in the aft cabin area as a supplemental oremergency source of oxygen. The mask is stowed behindan overhead cover placarded FIRST AID OXYGEN PULL.Removing the cover allows the mask to drop out of thecontainer, exposing a manual control valve, whichreleases oxygen to the mask when placed in the ONposition. After using the mask, the manual valve in thecontainer must be turned OFF before stowing the maskand replacing the cover.

g. Oxygen Duration Example Problem.WANTED Duration in minutes of oxygen

at 100% capacity.KNOWN Two man crew plus one passen-

ger, cabin pressure altitude =15,000 feet, crew masks, normal,100% capacity.

METHOD Find "two man crew plus onepass" line, move right then downto 15,000 - "normal" read "232.1" minutes.

WANTED Duration of oxygen for previousexample data at 84% of capacity.

KNOWN 232.1 minutes duration at 100%,84% capacity, total aircraft flow= 13.9 LPM.

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METHOD Multiply 232.1 X 0.84 = 194.9minutes. or Multiply 3,226 X0.84 = 2709.8, divide by 13.9LPM = 194.9 minutes.

WANTED Duration of oxygen for comple-ment at other cabin pressure al-titude, at less than 100% capaci-ty.

KNOWN Cylinder at 84% capacity, 100%capacity = 3,226 L, cabin pres-sure altitude = 21,000 feet.I crew mask = 7.2 LPM (100%),I passenger mask = 3.7 LPM

METHOD Multiply 3,226 L X 0.84 = 2.709.8 L, multiply 2 crew X 7.2LPM = 14.4 LPM, multiply 1passenger X 3.7 LPM, add 14.4LPM crew plus 3.7 LPM passen-ger = 18.1 LPM.Divide 3,226 Lby 18.1 LPM = 178.2 minutes.

h. Oxygen Cylinder Capacity Example Problem.WANTED

a. Percent of capacity at known pressure andtemperature.

b. Pressure when temperature decreases.

KNOWN Pressure = 1,600 PSIG stabi-lized cylinder temperature is es-timated at 20° C decreased stabi-lized cylinder temperature is es-timated at -30° C.METHOD

c. Enter 1600 PSIG move up to 20° C line, moveright to 84%.

d. Move left on 84% line to -30° C line, move downto 1250 PSIG.

WANTED 100% capacity pressure atknown temperature.

KNOWN Temperature -- -30° C.METHOD Move left along 100% line to

-30° C line and move down to1420 PSIG.

2-62. WINDSHIELD WIPERS.a. Description. Two electrically operated windshield

wipers, are provided for use at takeoff, cruise and landingspeed. A rotary switch (fig. 2-18) placardedWINDSHIELD WIPER, located on the overhead controlpanel, selects mode of windshield wiper operation. Aninformation placard above the switch states: DO NOTOPERATE ON DRY GLASS. Function positions on theswitch, as read clockwise, are placarded: PARK OFFSLOW FAST. When the switch is held in the spring-loaded PARK setting, the blades will return to their normalinoperative position on the glass, then, when released, theswitch will return to OFF position terminating windshieldwiper operation. The FAST and SLOW switch positionsare separate operating speed

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Table 2-5. Oxygen Duration in Minutes 128 Cubic Foot System

TM 55-1510-219-10

Figure 2-20. Cylinder Capacity vs Pressure and Temperature

settings for wiper operation. The windshield wiper circuitis protected by one 10-ampere circuit breaker, placardedWSHLD WIPER, located on the overhead circuit breakerpanel (fig. 2-26).

CAUTIONDo not operate windshield wiperson dry glass. Such action candamage the linkage as well asscratch the windshield glass.

b. Normal Operation. To start, turn WINDSHIELDWIPER switch to FAST or SLOW speed, as desired. Tostop, turn the switch to the PARK position and release.The blades will return to their normal. inoperative positionand stop. Turning the switch only to the OFF position willstop the windshield wipers, without returning them to thenormal inactive position.

2-63. CIGARETTE LIGHTERS AND ASH TRAYS.

The pilot and copilot have individual cigarettelighters and ash trays mounted in escutcheons outboardof their seats. The cigarette lighters are protected by a 5-ampere circuit breaker, placarded CIGAR LIGHTER, onthe overhead circuit breaker panel (fig. 2-26).

2-64. ELECTRIC TOILET.

a. Description. An electric toilet is installed in the aftcabin area. The circuit is protected by a 10ampere circuitbreaker located in the power distribution panel under thefloor ahead of the main spar.

b. Operation. A switch, placarded PRESS TOFLUSH, is mounted on the seat assembly for operation ofthe toilet. Pressing the switch applies DC power to themotor which drives the pump. The pump applies flushingfluid through a nozzle in the upper rim and washes theinner surface of the bowl. Waste is carried to the wastetank mounted below the bowl. When desired, theremovable waste tank may be removed from the toilet forservicing.

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2-65. SUN VISORS.

CAUTIONWhen adjusting the sun visors,grasp only by the top metalattachment to avoid damage tothe plastic shield.

A sun visor is provided for the pilot and copilotrespectively (fig. 2-9). Each visor is manually adjust

able. When not needed as a sun shield, each visor maybe manually rotated to a position flush with the top ofthe cockpit so that it does not obstruct view through thewindows.

2-66. RELIEF TUBES.

One relief tube is provided, located immediately aft ofthe cabin door, on the left side of the fuselage.

Section VIII. HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL CONTROL SYSTEM

2-67. HEATING SYSTEM.

a. Description. Warm air for heating the cockpitand mission avionics compartments and warmwindshield defrosting air is provided by bleed air fromboth engines (fig. 2-21). Engine bleed air is combinedwith ambient air in the heating and pressurization flowcontrol unit in each nacelle. If the mixed bleed air is toowarm for cockpit comfort, it is cooled by being routedthrough an air-to-air heat exchanger located in theforward portion of each inboard wing. If the mixed bleedair is not too warm, the air-to-air heat exchangers arebypassed. The mixed bleed air is then ducted to amixing plenum, where it is mixed with cabin recirculatedair. The warm air is then ducted to the cockpit outlets,windshield defroster outlets, and to the floor outlets inthe mission avionics compartment.

(1) Bleed airflow control unit. A bleed air flowcontrol unit, located forward of the firewall in eachengine nacelle controls the flow of bleed air and themixing of ambient air to make up the total airflow to thecabin for heating, windshield defrosting, pressurizationand ventilation. The unit is fully pneumatic except foran integral electric solenoid firewall shutoff valve,controlled by the bleed air switches located on theoverhead control panel (fig. 2-18) and a normally opensolenoid valve operated by the left landing gear safetyswitch.

(2) Pneumatic bleed air shutoff valve. Apneumatic shutoff valve is provided in each nacelle tocontrol the flow of bleed air to the surface, antenna andbrake device systems. These valves are controlled bythe bleed air valve switches located on the overheadcontrol panel (fig. 2-18).

(3) Bleed air valve switches. The bleed airflow control unit shutoff valve and pneumatic bleed airshutoff valves are controlled by two switches placardedBLEED AIR VALVE OPEN ENVIRO OFF -PNEU &ENVIRO OFF, located on the overhead control panel(fig, 2-18). When set to the open position, both the

environmental flow control unit shutoff valve and thepneumatic shutoff valve are open; when set to theENVIRO OFF position, the environmental flow controlunit shutoff valve is closed, and the pneumatic bleed airvalve is open; in the PNEU & ENVIRO OFF position,both are closed. For maximum cooling on the ground,turn the bleed air valve switches to the ENVIRO OFFposition.

(4) Cabin temperature mode selector switch.A switch placarded CABIN TEMP MODE MAN COOLMAN HEAT OFF AUTO A/C COLD OPN, located on theoverhead control panel, controls cockpit and missionavionics compartment heating and air conditioning.When the cabin temperature mode selector switch is setto the AUTO position, the heating and air conditioningsystems are automatically controlled. Control signalsfrom the temperature control box are transmitted to thebleed air heat exchanger bypass valves. Here thetemperature of the air flowing to the cabin is regulatedby the bypass valves controlling the amount of airbypassing the heat exchangers. When the cabintemperature mode selector switch is set to the AUTOposition, the heating and air conditioning systems areautomatically controlled. When the temperature of thecabin has reached the temperature setting of the cabintemperature control rheostat, the automatic temperaturecontrol allows hot air to bypass the air-to-air exchangers.When both bypass valves are in the fully closedposition, allowing no air to bypass the heat exchangers,the air conditioner begins to operate, providingadditional cooling. When the cabin temperature modeselector switch is set to the A/C COLD OPN position, theair conditioning system is in continuous operation. Thecabin temperature control rheostat, in conjunction withthe cabin temperature control sensor, providesregulation of cockpit and mission equipmentcompartment temperature. Bleed air heat is added asre

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Figure 2-21. Environmental Systems

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quired to maintain the temperature selected by the cabintemperature control rheostat.

(5) Cabin temperature control rheostat. Acontrol knob placarded CABIN TEMP INCR, located onthe overhead control panel (fig. 2-18), providesregulation of cabin temperature when the cabintemperature mode selector switch is set to the AUTOposition, or the A/C COLD OPN position. A etemperature sensing unit in the cabin, in conjunctionwith the setting of the cabin temperature controlrheostat, initiates a heat or cool command to thetemperature controller for the desired cockpit or missionavionics compartment environment.

(6) Manual temperature control switch. Aswitch placarded MANUAL TEMP INCR DECR, locatedon the overhead control panel (fig. 2-18), controlscockpit and mission avionics compartment temperaturewith the cabin temperature mode selector switch set tothe MAN HEAT positions. The manual temperaturecontrol switch controls cockpit and mission avionicstemperature by providing a means of manually changingthe amount that the bleed air bypass valves are openedor closed. To increase cabin temperature the switch isheld to the INCR position. To decrease cabintemperature, the switch is held to the DECR position.Approximately 30 seconds per valve is required to drivethe bypass valves to the fully open or fully closedposition. Only one valve moves at a time.

(7) Forward vent blower switch. The forwardvent blower is controlled by a switch placarded VENTBLOWER AUTO LO HI, located on the overhead controlpanel (fig. 2-18). In the auto position the fan will run atlow speed except when the cabin temperature modeselector switch is set to the OFF position, in this casethe blower will not operate.

(8) Aft vent blower switch. The aft ventblower is controlled by a switch placarded AFT VENTBLOWER OFF AUTO ON, located on the overheadcontrol panel (fig. 2-18). The single speed bloweroperates automatically through the cabin temperaturemode selector switch when the aft vent blower switch isplaced in the AUTO position. The blower runscontinuously when the switch is placed in the ONposition, In the OFF position, the blower will not operate.

b. Automatic Heating Mode.

1. Bleed air valve switches OPEN,LEFT and RIGHT.

2. Cabin temperature mode selectorswitch AUTO.

3. Cabin temperature control rheostatAs required.

4. Cabin, cockpit and defrost air knobsAs required

c. Cabin Heating Mode.1. Bleed air valve switches OPEN,

LEFT and RIGHT.

2. Cabin temperature mode selectorswitch MAN HEAT.

3. Vent blower switches As required.

4. Manual temperature switch Asrequired.

5. Cabin, cockpit and defrost air knobsAs required.

2-68. AIR CONDITIONING SYSTEM.

a. Description. Cabin air conditioning is providedby a refrigerant gas vapor cycle refrigeration systemconsisting of a belt driven, engine mounted compressor,installed on the #2 engine accessory pad, refrigerantplumbing, N1 speed switch, high and low pressureprotection switches, condenser coil, condenser blower,forward and aft evaporator, receiver dryer, expansionvalve and a bypass valve. The plumbing from thecompressor is routed through the right inboard wingleading edge to the fuselage and then forward to thecondenser coil, receiver-dryer, expansion valve, bypassvalve, and forward evaporator, which are located in thenose of the aircraft. A circuit breaker placarded AIRCOND CONTR, located on the overhead circuit breakerpanel (fig. 226), protects the compressor clutch circuit.

(1) Forward evaporator. The forwardevaporator and blower supplies the cockpit, forwardceiling outlets, and forward floor outlets. The forwardevaporator blower has a high speed which can beselected by setting the VENT BLOWER switch, locatedon the overhead control panel (fig. 2-18), to the HIposition. The forward vent blower is protected by acircuit breaker located on the DC power distributionpanel, located in the forward equipment bay.

(2) Aft evaporator. The aft evaporator andblower are located in the fuselage center aisleequipment bay aft of the rear spar. Environmental air iscirculated through the evaporator in either manual orautomatic control mode. The rear evaporator suppliesthe aft ceiling outlets, rear floor outlets, and toiletcompartment. Rear evaporator blower is protected by acircuit breaker located on the DC power distributionpanel in the lower equipment bay.

(3) High and low pressure limit switches.High and low pressure limit switches are provided toprevent compressor operation beyond operational

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limits. When the low or high pressure switches areactivated, a reset switch/light assembly located in thenosewheel well is activated to prevent furthercompressor operation.

(4) Thermal sense switch. A 33° thermalsense switch is installed on the forward evaporator.This sense switch actuates a hot gas bypass valvewhich bleeds off a portion of the refrigerant from theforward evaporator, thereby preventing icing of theevaporator.

(5) Condenser blower. A vane-axial blowerdraws air through the condenser on the ground as wellas in flight. The current limiter for this blower is locatedon the DC distribution panel in the lower equipment bay.When the cabin temperature mode selector switch is setto the A/C COLD OPN position, the condenser blowerwill be off, and will remain off until the added highpressure switch senses a compressor dischargepressure equal to the pressure it is set to. Thecondenser blower will then remain in operation until thelow pressure switch senses that the system pressure hasdropped to the pressure it is set to.

(6) Air conditioning cold start bypass valve.Low ambient temperature operation of the airconditioning system is accomplished by means of abypass valve located in the nose wheel well, anadditional high pressure switch, and an additional lowpressure switch, both of which are located in the rightinboard wing leading edge. The cold start bypass valveopens when the cabin temperature mode selector switchis set to the A/C COLD OPN position to enhancerefrigerant flow at low temperatures and improve airconditioner performance.

b. Normal Operation.

(1) Automatic cooling mode.


2. Cabin temperature mode selectorswitch AUTO.



(2) Manual cooling mode.


NOTE

For maximum cooling on the ground,set the bleed air valve switches to theENVIRO OFF position.

2. Cabin temperature mode selectorswitch MAN COOL.

(3) Air conditioning cold start mode. (Used ifambient temperature is between 10°C and -25°C).

NOTE

Setting the cabin temperaturemode selector switch to the A/CCOLD OPN position at ambienttemperatures below -25°C maycause the air conditioner systemto exceed the compressor lowpressure limit switch setting,illuminating the reset switch inthe nosewheel well, therebyrendering the system inoperativefor the remainder of the flight.


2. Cabin temperature mode selectorswitch A/C COLD OPN.



2-69. UNPRESSURIZED VENTILATION.

Ventilation is provided by two sources. Onesource is through the bleed air heating system in boththe pressurized and unpressurized mode. The secondsource of ventilation is obtained from ram air throughthe condenser section in the nose through a check valvein the vent blower plenum. Ventilation from this sourceis in the unpressurized mode only with the CABINPRESS switch in the DUMP position. The check valvecloses during pressurized operation. Ram air ventilationis distributed through the main ducting system to alloutlets. Ventilation air, ducted to each individual eyeballcold air outlet, can be directionally controlled by movingthe ball in the socket. Volume is regulated by twistingthe outlet to open or close the valve.

2-70. ENVIRONMENTAL CONTROLS.

An environmental control section on the overheadcontrol panel (fig. 2-18) provides for automatic ormanual control of the system. This section contains allthe major controls of the environmental functionincluding bleed air valve switches, a vent blower controlswitch, an aft vent blower switch, a manual temperatureswitch for control of the heat exchanger valves, a cabintemperature level control,

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and the cabin temp mode selector switch for selectingautomatic heating or cooling or manual heating orcooling. Four additional manual controls on the maininstrument subpanels may be utilized for partialregulation of cockpit comfort when the cockpit curtain isclosed and the cabin comfort level is satisfactory.

a. Heating Mode.

(1) If the cockpit is too cold:1. Pilot and copilot air knobs As

required.

2. Defrost air knob As required.

3. Cabin air knob Pull out in smallincrements. Allow 3 5 minutes aftereach adjustment for system tostabilize.

(2) If the cockpit is too hot:1. Cabin air knob As required.

2. Pilot and copilot air knobs In asrequired.

3. Defrost air knob In as required.

b. Cooling Mode.

(1) If the cockpit is too cold:1. Pilot and copilot air knob In as

required.

2. Defrost air knob In as required.

3. Overhead cockpit outlets Asrequired.

(2) If the cockpit is too hot:1. Pilot and copilot air knobs Out as

required.

2. Cabin air knob. Close in smallincrements. Allow 3 5 minutes aftereach adjustment for system tostabilize. If CABIN AIR knob iscompletely closed before obtainingsatisfactory cockpit comfort, it maybe necessary to place the aft ventblower switch in the ON position toactivate the aft evaporator torecirculate cabin air.

c. Automatic Mode Control. When the AUTOmode is selected on the cabin temperature modeselector switch, the heating and air conditioning systemsare automatically controlled. When the temperature ofthe cabin has reached the selected setting, theautomatic temperature control allows heated air tobypass the air-to-air exchangers in the wing center TM

section. The warm bleed air is mixed with the cooledair. The rear evaporator picks up recirculated cabin aironly.

(1) When the automatic control drives theenvironmental system from a heat mode to a coolingmode, the bypass valves close. When the left bypassvalve reaches a fully closed position, the refrigerationsystem provided the right engine N1 speed is above62°. When the bypass valve is opened to a positionapproximately 30° from full closed, the refrigerationsystem will turn off.

(2) The CABIN TEMP INCR control providesregulation of the temperature level in the automaticmode. A temperature sensing unit in the cabin, inconjunction with the control setting, initiates a heat orcool command to the temperature controller for desiredcockpit and cabin environment.

d. Manual Mode Control. With the cabintemperature mode selector in the MAN HEAT or MANCOOL position, regulation of the cabin temperature isaccomplished manually with the MANUAL TEMP switch.

(1) In the MAN HEAT mode, the automaticsystem is overridden and the system is controlled byopening and closing the bypass valves (two) with theMANUAL TEMP INCR DECR switch. To increase cabintemperature, hold the switch at the INCR position, todecrease cabin temperature, hold the switch in theDECR position. Allow approximately 30 seconds pervalve to drive the bypass valves to the fully open or fullyclosed position. Only one valve moves at a time.

(2) With the cabin temperature selectorswitch in the MAN COOL position, the automatictemperature control system is bypassed. In the manualcooling mode, the refrigeration system is on, providingthe right engine right engine N1speed is above 62%,however, the bypass valves may be manually positionedfor the desired temperature. Hold the MANUAL TEMPswitch in the DECR position approximately one minuteto fully close air-to-air heat exchanger bypass valves.

e. Bleed Air and Vent Control.

(1) Bleed air entering the cabin is controlledby bleed air valve switches placarded BLEED AIRVALVE OPEN ENVIRO OFF PNUE & ENVIRO OFF.When the switch is in the OPEN position, theenvironmental flow control unit and the pneumatic valveare open. When the switch is in the ENVIRO OFFposition, the environmental flow control unit is closedand the pneumatic bleed air valve is open; in the PNEU& ENVIRO OFF position, both are closed. Formaximum cooling on the

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ground, turn the bleed air valve switches to the ENVIROOFF position.

(2) The forward vent blower is controlled by aswitch placarded VENT BLOWER AUTO LOW HI. TheHI and LOW positions regulate the blower to two speedsof operation. IN the AUTO position, the fan will run at

low speed except when the CABIN TEMP mode selectorswitch is placed in the OFF position. In the OFFposition, the blower will not operate.

(3) The aft vent blower is controlled by aswitch placarded AFT VENT BLOWER OFF AUTO ON.In the OFF position, the blower will not operate.

Section IX. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM

2-71. DESCRIPTION.

The aircraft employs both direct current (DC) andalternating current (AC) electrical power. The DCelectrical supply is the basic power system energizingmost aircraft circuits. Electrical power is used to startthe engines, to power the landing gear and flap motors,and to operate the standby fuel pumps, ventilationblower, lights and electronic equipment. AC power isobtained from DC power through inverters. The threesources of DC power consist of one 20 cell 34-ampere/hour battery and two 400ampere starter-generators. DC power may be applied to the aircraftthrough an external power receptacle on the rightnacelle. The starter-generators are controlled bygenerator control units. The output of each generatorpasses through a cable to the respective generator bus(fig. 2-22). Other buses distribute por to aircraft DCloads, and derive power from the generator buses. Thegenerators are paralleled to balance the DC loadsbetween the two units. When one of the generators isnot on line, aircraft DC power requirements continue tobe supplied from one of the other generators. Most DCdistribution buses are connected to both generatorbuses but have isolation diodes to prevent powercrossfeed between the generating systems, whenconnection between the generator buses is lost. Thus,when either generator is lost because of a ground fault,the operating generator will supply power for all aircraftDC loads except those receiving power from theinoperative generator's bus which cannot be crossfed.When a generator is not operating, reverse current andover-voltage protection is automatically provided. Twoinverters operating from DC power produce the requiredsingle-phase AC power (figure 2-23). Three phase ACelectrical power for inertial navigation system andmission avionics is supplied by two DC poweredinverters (figures 2-24 and 2-25).

2-72. DC POWER SUPPLY.

One nickel-cadmium battery furnishes DC powerwhen the engines are not operating. This 24-volt, 34-ampere/hour battery, located in the right wing centersection, is accessible through a panel on the 2-58 top ofthe wing. DC power is produced by two engine-driven28 volt, 400-ampere starter-generators. Controls and

indicators associated with the DC supply system arelocated on the overhead control panel (fig. 2-18) andconsists of a single battery switch (BATT), two generatorswitches (#1 GEN and #2 GEN), and two volt-loadmeters.

a. Battery Switch. A switch, placarded BATT islocated on the overhead control panel (fig. 2-18) underthe MASTER SWITCH. The BATT switch controls DCpower to the aircraft bus system through the batteryrelay, and must be ON to allow external power to enteraircraft circuits. When the MASTER SWITCH is movedaft, the BATT switch is forced OFF.

NOTE

With battery or external powerremoved from the aircraftelectrical system, due to fault,power cannot be restored to thesystem until the BATT switch ismoved to OFF/RESET, then ON.

b. Generator Switches. Two switches (fig. 218),placarded #1 GEN and #2 GEN are located on theoverhead control panel under the MASTER SWITCH.The toggle switches control electrical power from thedesignated generator to paralleling circuits and the busdistribution system. Switch positions are placardedRESET, ON and OFF. RESET is forward (spring-loaded back to ON), ON is center, and OFF is aft.When a generator is removed from the aircraft electricalsystem, due either to fault or from placing the GENswitch in the OFF position, the affected unit cannot haveits output restored to aircraft use until the GEN switch ismoved to RESET, then ON.

c. Master Switch. All electrical current may beshut off using the MASTER SWITCH gangbar (fig.

2-18) which extends below the battery and generatorswitches. The MASTER SWITCH gangbar is movedforward when a battery or generator switch is turned on.When moved aft, the bar forces each switch to the OFFposition.

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Figure 2-22. DC Electrical System (Sheet 1 of 3)

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Figure 2-23. Single Phase AC Electrical System

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Figure 2-24. Three Phase AC Electrical System

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Figure 2-25. Mission Equipment Power System

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Figure 2-26. Overhead Circuit Breaker Panel

d. Volt-Loadmeters. Two meters (fig. 2-18), onthe overhead control panel display voltage readings andshow the rate of current usage from left and rightgenerating systems. Each meter is equipped with aspring-loaded push-button switch which when manuallypressed will cause the meter to indicate main busvoltage. Each meter normally shows output amperagereading from the respective generator, unless the push-button switch is pressed to obtain bus voltage reading.Current consumption is indicated as a percentage oftotal output amperage capacity for the generatingsystem monitored.

e. Battery Monitor. Nickel-cadmium batteryoverheating will cause the battery charge current toincrease if thermal runaway is imminent. The aircrafthas a charge-current sensor which will detect a chargecurrent. The charge current system senses batterycurrent through a shunt in the negative lead of thebattery. Any time the battery charging current exceedsapproximately 7-amperes for 6 seconds or longer, theyellow BATTERY CHARGE annunciator light and themaster fault caution light will illuminate. Following abattery engine start, the caution light will illuminateapproximately six seconds after the generator switch isplaced in the ON position. The light will normally

extinguish within two to five minutes, indicating that thebattery is approaching a full charge. The time intervalwill increase if the battery has a low state of charge, thebattery temperature is very low, or if the battery haspreviously been discharged at a very low rate (i.e.,battery operation of radios or lights for prolongedperiods). The caution light may also illuminate for shortintervals after landing gear and/or flap operation. If thecaution light should illuminate during normal steady-state cruise, it indicates that conditions exist that maycause a battery thermal runaway. If this occurs, thebattery switch shall be turned OFF and may be turnedback ON only for gear and flap extension and approachto landing. Battery may be used after a 15 to 20 minutecool down period.

f. Generator Out Warning Lights. Twocaution/advisory annunciator panel lights inform the pilotwhen either generator is not delivering current to theaircraft DC bus system. These lights are placarded #1DC GEN and #2 DC GEN. Two MASTER CAUTIONlights and illumination of either fault light indicates thateither the identified generator has failed or voltage is notsufficient to keep it connected to the power distributionsystem.

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CAUTION

The GPU shall be adjusted toregulate at 28 volts maximum toprevent damage to the aircraft.

g. DC External Power Source. External DC powercan be applied to the aircraft through an external powerreceptacle on the underside of the right wing leadingedge just outboard of the engine nacelle. Thereceptacle is installed inside of the wing structure and isaccessible through a hinged access panel. DC power issupplied through the DC external plug and applieddirectly to the battery bus after passing through theexternal power relay. Turn off all external power whileconnecting the power cable to, or removing it from, theexternal power supply receptacle. The holding coilcircuit of the relay is energized by the external powersource when the BATT switch is in the ON position. TheGPU shall be adjusted to regulate at 28 volts maximumto prevent damage to the aircraft battery.

h. Security Keylock Switch. The aircraft has asecurity keylock switch (fig. 2-18) installed on theoverhead control panel, placarded OFF ON. The switchis connected to the battery relay circuit and must be ONwhen energizing the battery master power switch. Thekey cannot be removed from the lock when in the ONposition.

i. Circuit Breakers. The overhead circuit breakerpanel (fig. 2-26) contains circuit breakers for mostaircraft systems. The circuit breakers on the panel aregrouped into areas which are placarded as to thegeneral function they protect. A DC power distributionpanel is mounted beneath the floor forward of the mainspar. This panel contains higher current rated circuitbreakers and is not accessible to the flight crew undernormal conditions.

2-73. AC POWER SUPPLY.

a. Single Phase AC Power Supply. AC power forthe aircraft is supplied by inverter units, numbered #1and #2 (fig. 2-23) which obtain operating current fromthe DC power system. Both inverters are rated at 750volt-amperes and provide single phase output only.Each inverter provides 115 volt and 26 volt 400 Hz ACoutput. The inverters are protected by circuit breakersmounted on the DC power distribution panel mountedbeneath the floor. Controls and indicators of the ACpower system are located on the overhead control paneland on the caution/advisory annunciator panel.

b. AC Power WARNING/CAUTION Lights. TwoMASTER/CAUTION lights and the illumination of anannunciator caution light #1 INVERTER or #2INVERTER indicates an inverter failure.

c. Instrument AC Light. A red warning light,located on the warning annunciator panel, placardedINST AC, will illuminate if all instrument AC bussesshould fail.

d. Inverter Control Switches. Two switches (fig.2-18), placarded INVERTER #1 and #2 on the overheadcontrol panel give the pilot control of inverters single-phase AC power.

e. Volt-Frequency Meters. Two volt-frequencymeters (fig. 2-18) are mounted in the overhead controlpanel to provide monitoring capability for both 115 VACbuses. Normal display on the meter is shown infrequency (Hz). To read voltage, press the buttonlocated in the lower left corner of the meter. Normaloutput of the inverters will be indicated by 115 VAC and400 Hz on the meters.

f Three Phase AC Power Supply. Three phaseAC electrical power for operation of the inertialnavigation system and mission avionics is supplied byeither of two DC powered 3000 volt-ampere solid statethree phase inverters (fig. 2-24).

(1) Three phase inverter control switch. Athree position switch placarded #1 MISSION INV, OFF,#2 MISSION INV, located on the mission control panelcontrols three phase inverter operation.

(2) Three phase volt-frequency meter. Athree phase volt/frequency meter, mounted on themission control panel, monitors output of the selectedthree phase inverter. Frequency (Hz) is normallydisplayed on the meter. To read voltage press buttonlocated in the lower left corner of the meter.

(3) Three phase loadmeter. A three phaseloadmeter, mounted on the mission control panel,monitors inverter output level.

(4) Three phase AC off annunciator light. Anindicator light placarded 3( AC OFF, located on themission annunciator panel indicates that three phase ACpower is not being supplied.

(5) Three phase AC external power. Externalthree phase AC power for operation of the inertialnavigation system or mission equipment, can be appliedto the aircraft through an external power receptaclelocated on the underside of the left wing leading edgejust outboard of the engine nacelle. The receptacle isinstalled inside the wing structure and is accessiblethrough a hinged access panel. The AC electricalsystem is automatically isolated from the external powersource if the external power is over or under voltage,over or under frequency, or has an improper phasesequence.

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(a) External AC power annunciatorlight. An annunciator light placarded EXT AC PWR ON,located on the mission annunciator panel indicates thatan AC GPU plug is mated to the AC external powerreceptacle and the External AC Power control switch isOn.

(b) External AC power control switch.A switch placarded AC EXT POWER, located on themission control panel controls application of three phaseAC power to the aircraft.

Section X. LIGHTING

2-74. EXTERIOR LIGHTING.

a. Description. Exterior lighting (fig. 2-27)consists of a navigation light on top of the aft section ofthe vertical stabilizer, one navigation light on top andbottom of each wing tip pod, two strobe beacons, one ontop of the vertical stabilizer and one on the underside ofthe fuselage center section, dual landing lights, one taxilight mounted on the nose gear assembly, a recognitionlight located in each wing tip, and two ice lights, onelight flush mounted in each nacelle, positioned toilluminate along the leading edge of each outboard wing.

b. Navigation Lights. The navigation lights areprotected by a 5-ampere circuit breaker placarded NAVon the overhead circuit breaker panel (fig. 2-26).Control of the lights is provided by a switch placardedNAV-ON on the overhead control panel (fig. 2-18).

c. Strobe Beacons. The strobe beacons are dualintensity units. They are protected by a 15amperecircuit breaker placarded BCN on the overhead circuitbreaker panel (fig. 2-26). Control of the lights isprovided by a switch placarded BEACONDAY-NIGHT(fig. 2-18). Placing the switch in the DAY position willactivate the high intensity white section of the strobelights for greater visibility during daytime operation.Placing the switch in the NIGHT position activates thelower intensity red section of the strobe lights.

d. Landing/Taxi Lights. Dual landing lights and asingle taxi light are mounted on the nose gear assembly.The lights are controlled by switches, placardedLANDING and TAXI, located in the LIGHTS section ofthe pilot's subpanel. In addition, these light areextinguished whenever the landing gear is retracted.The landing/taxi lights circuits are protected by 5-ampere circuit breakers placarded LANDING and TAXIrespectively, located on the overhead circuit breakerpanel (fig. 2-26). Additional circuit protection isprovided by 35-ampere and 15ampere circuit breakerson the DC power distribution panel located beneath thecockpit floor.

e. Ice Lights. The ice lights circuit is protected bya 5-ampere circuit breaker placarded ICE on theoverhead circuit breaker panel (fig. 2-26). Control of

the lights is provided by a switch placarded ICE on theoverhead control panel (fig. 2-18). Prolonged useduring ground operation may generate enough heat todamage the lens.

f. Recognition Lights. A RECOG switch, locatedin the pilot's subpanel LIGHTS section, controls a whiterecognition light in each wing tip. When requested, thissteady, bright light is used for identification in the trafficpattern. The recognition lights circuit is protected by a 71/2 ampere RECOG circuit breaker located on theoverhead circuit breaker panel (fig. 2-26).

2-75. INTERIOR LIGHTING.

Lighting systems are installed for use by the pilotand copilot and by the passengers in the cabin area.The lighting systems in the cockpit are provided withintensity controls on the overhead control panel. Aswitch placarded MASTER PANEL LIGHTS on theoverhead control panel (fig. 2-18) provides overall on-off control for all engine instrument lights, pilot andcopilot instrument lights, overhead panel lights, consoleand subpanel lights and the free air temperature light.

a. Cockpit Lighting.

(1) Flight instrument lights. Each individualflight instrument contains internal lamps for illumination.The circuit is protected by a 7 1/2-ampere circuitbreaker placarded FLT INST on the overhead circuitoverhead circuit breaker panel (fig. 226). Control isprovided by two rheostat switches placarded PILOT orCOPILOT INST LIGHTS-OFF-BRT on the overheadcontrol panel (fig. 2-18). Turning the control clockwisefrom OFF turns the lights on and increases theirbrilliance.

(2) Instrument indirect lights. Three lights aremounted in the glareshield overhang along the top edgeof the instrument panel and provide overall instrumentpanel illumination. The circuit is protected by a 5-ampere circuit breaker placarded INST INDIRECT onthe overhead circuit breaker panel (fig. 2-26). Controlis provided by a rheostat switch placarded INSTINDIRECT LIGHTS-OFF-BRT on the overhead controlpanel (fig. 2-18). Turning the

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Figure 2-27. Exterior Lighting

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control clockwise from OFF turns the lights on andincreases their brilliance.

(3) Engine instrument lights. Each individualengine instrument contains internal lamps forillumination. The circuit is protected by a 7 1/2amperecircuit breaker placarded FLT INST on the overheadcircuit breaker panel (fig. 2-26). Control is provided bya rheostat switch placarded ENGINE INST LIGHTS OFFBRT on the overhead control panel (fig. 2-18). Turningthe control clockwise from OFF turns the lights on andincreases their brilliance.

(4) Flood light. A single overhead flood lightis installed. It provides overall illumination of the entirecockpit area. The circuit is protected by a 5-amperecircuit breaker mounted beneath the battery andconnected to the emergency battery bus. Control isprovided by a rheostat switch placarded OVERHEADFLOODLIGHT-OFF-BRT on the overhead control panel(fig. 2-18). Turning the control clockwise from OFFturns the light on and increases its brilliance.

(5) Overhead panel lights. Lamps on theoverhead circuit breaker panel, control panel, and fuelmanagement panel are protected by a 7 1/2amperecircuit breaker placarded LIGHTS OVHD on theoverhead circuit breaker panel (fig. 2-26). Control isprovided by a rheostat switch placarded OVERHEADPANEL LIGHTS-OFF-BRT on the overhead controlpanel (fig. 2-18). Turning the control clockwise from offturns the lights on and increases their brilliance.

(6) Subpanel and console lights. Lights onthe pilot's and copilot's subpanels, console edge litpanels and pedestal extension panels are protected by a7 1/2-ampere circuit breaker placarded LIGHTSSUBPANEL & CONSOLE on the overhead circuitbreaker panel (fig. 2-26). Control is provided by tworheostat switches placarded SUBPANEL or CONSOLELIGHTS-OFF-BRT on the overhead control panel (fig.2-18). Turning the control clockwise from OFF turns thelights on and increases their brilliance.

(7) Free air temperature light. Two post lightsare mounted adjacent to the free air temperature gageon the left cockpit sidewall trim panel. The circuit isprotected by a 7 1/2-ampere circuit breaker placardedFLT INST on the overhead circuit breaker panel (fig. 2-26). Control is provided by a push button switchadjacent to the gage. No intensity control is provided.

b. Cabin Lighting.

(1) Interior lights. Three cabin lights areinstalled in the overhead trim. The circuit is protected

by a 10-ampere circuit breaker placarded CABIN on theoverhead circuit breaker panel (fig. 2-26). Control isprovided by a switch placarded CABIN LIGHTS-BRIGHT-DIM-OFF on the subpanel (fig. 2-6).

(2) Threshold and spar cover lights. A sparcover light is installed on the left side of the sunken aisleimmediately aft of the main spar cover. Both circuitsare protected by a 5-ampere circuit breaker mountedbeneath the battery and connected to the emergencybattery bus. Both lights are controlled by the switchmounted forward of the airstair door. If the lights areilluminated, closing the cabin door will automaticallyextinguish them.

(3) Dome light. A dome light is installed inthe baggage area, in the overhead. The circuit isprotected by a 5-ampere circuit breaker mountedbeneath the battery and connected to the emergencybattery bus. Control is provided by a switch mountedadjacent to the light.

(4) Cabin utility light. There is a cabin utilitylight adjacent to each cabin light. Each utility light isindividually controlled by a rheostat placarded OFF-ON-BRT on the back of the light. There is a momentary ONswitch in the center of the rheostat. Each light iscapable of producing a red or white spotlight by turningthe selector on the front of the light. To remove the lightfrom the stationary position, loosen the retaining screwdirectly below the light escutcheon and pull down on thelight. The light is connected to the light housing by an11 inch coiled cord that extends to approximately 50inches. The CABIN LIGHTS switch must be on for utilitylamps to operate.

2-76. EMERGENCY LIGHTING.

a. Description. An independent battery operatedlighting system is installed. The system is actuatedautomatically by shock, such as a forced landing. Itprovides adequate lighting inside and outside thefuselage to permit the crew to read instruction placardsand locate exits. An inertia switch, when subjected to a2 3 G shock, will illuminate interior lights in the cockpit,forward and aft cabin areas, and exterior lights aft of theemergency exit and aft of the cabin door. The batterypower source is automatically recharged by the aircraftelectrical system.

b. Operation. An emergency lights overrideswitch, located on the overhead control panel (fig. 218),is provided to turn the system off if it is accidentallyactuated. The switch is placarded EMERG LIGHTSOVRD OFF-RESET-AUTO-TEST. Should the systemaccidentally actuate, placing the switch in themomentary OFF RESET position will extinguish

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the lights. To test the system, place the switch in themomentary TEST position. The lights should illuminate.

Moving the switch to the OFF-RESET position will turnthe system off and reset it.

Section XI. FLIGHT INSTRUMENTS

2-77. TURN-AND-SLIP INDICATORS.

Turn-and-slip indicators are installed separatelyon the pilot and copilot sides of the instrument panel(fig. 2-28). The pilot's indicator provides yaw dampinginformation to the autopilot. These indicators aregyroscopically operated. They use DC power and areprotected by 5-ampere circuit breakers placarded TURN& SLIP PILOT or COPILOT on the overhead circuitbreaker panel (fig. 2-26).

2-78. AIRSPEED INDICATORS.

Airspeed indicators are installed separately on thepilot and copilot sides of the instrument panel (fig. 2-28). These indicators require no electrical power foroperation. The indicator dials are calibrated in knotsfrom 40 to 300. A striped pointer automatically displaysthe maximum allowable airspeed (245 KIAS, 0.47 mach)at the aircraft’s' present altitude.

2-79. PILOT'S ENCODING ALTIMETER.

The altimeter is located on the upper left side ofthe instrument panel (fig. 2-28). The altimeter is a self-contained unit which consists of a precision pressurealtimeter combined with an altitude encoder. Thedisplay indicates and the encoder transmits,simultaneously, pressure altitude information to thetransponder. Altitude is displayed on the altimeter by a10,000 foot counter, a 1000 foot counter, and a singleneedle pointer which indicates hundreds of feet on acircular scale in 50 foot increments. Below an altitudeof 10,000 feet, a diagonal warning symbol will appear onthe 10,000 foot counter. A barometric pressure settingknob is provided to insert the desired altimeter setting ininches Hg or millibars. A DC powered vibrator operatesinside the altimeter whenever aircraft power is on. If DCpower to the altitude encoder is lost, a warning flagplacarded CODE OFF will appear in the upper centerportion of the instrument face, indicating that the altitudeencoder is inoperative and that the system is notreporting altitude to ground stations. Operatinginstructions are contained in Chapter 3.

2-80. COPILOT'S ALTIMETER.

The copilot's altimeter is located on the upperright side of the instrument panel (fig. 2-28). It displaysaltitude by means of a 10,000 foot counter, a 1000 footcounter, a 100 foot counter, and a single needle pointerthat indicates on a circular scale marked in 50 footintervals. Below an altitude of 10, 000 feet, a diagonallystriped symbol covers the 10, 000 foot indicator. A knobis provided at the bottom right corner of the altimeter forsetting readings in the pressure windows.

2-81. VERTICAL VELOCITY INDICATORS.

Vertical velocity indicators are installed separatelyon the pilot and copilot sides of the instrument panel(fig. 2-28). They indicate the speed at which the aircraftascends or descends based on changes in atmosphericpressure. The indicator is a direct reading pressureinstrument requiring no electrical power for operation.

2-82. ACCELEROMETER.

The accelerometer, located on the instrumentpanel registers and records positive and negative Gloads imposed on the aircraft. One hand moves in thedirection of the G load being applied while the other two,one for positive G loads and one for negative G loads,follow the indicating pointer to its maximum travel. Therecording pointers remain at the respective maximumtravel positions of the G's being applied, providing arecord of maximum G loads encountered. Depressingthe push-to-reset knob at the lower left corner of theinstrument allows the recording pointers to return to thenormal position.

2-83. FREE AIR TEMPERATURE (FAT) GAGE.

The free air temperature gage, mounted outboardof the pilot's seat indicates the free air temperature indegrees Celsius.

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2-84. STANDBY MAGNETIC COMPASS.

WARNING

Inaccurate indications on thestandby magnetic compass willoccur while windshield heatand/or air conditioning is beingused.

The standby magnetic compass is located belowthe overhead fuel management panel to the right of thewindshield divider. It may be used in the event of failureof the compass system, or for instrument cross check.Readings should be taken only during level flight sinceerrors may be introduced by turning or acceleration. Acompass correction chart indicating deviation is locatedon the magnetic compass.

2-85. MISCELLANEOUS INSTRUMENTS.

a. Annunciator Panels. Three annunciator panelsare installed. One is a warning panel with red faultidentification lights, and the others are caution/ advisorypanels with yellow and green identification lights. thewarning panel is mounted near the center of theinstrument panel below the glareshield (fig. 228) andone caution/advisory panel is located on the centersubpanel. The mission annunciator panel is located onthe copilot's sidewall. Some normal flight operationsinvolve indications from the mission control panel,described in Chapter 4. Illumination of a red warninglight signifies the existence of a hazardous conditionrequiring immediate corrective action. A yellow cautionlight signifies a condition other than hazardous requiringpilot attention. A green advisory light indicates afunctional situation. Table 2-6 provides a list of causesfor illumination of the individual annunciator lights. Infrontal view both panels present rows of small, opaquerectangular indicator lights. Word printing on eachindicator identifies the monitored function, situation, orfault condition, but cannot be read until the light isilluminated. The bulbs of all annunciator panel lightsare tested by activating the ANNUNCIATOR TESTswitch, located on the right side of the caution/ advisorypanel. The system is protected by two 5ampere circuitbreakers placarded ANN PWR and ANN IND on theoverhead circuit breaker panel (fig. 2-26). Theannunciator system lights are dimmed when theMASTER PANEL LIGHTS switch is actuated and thepilot's flight instrument lights are on. The lights areautomatically reset to maximum brightness if:

(1) The main aircraft power (both DCgenerators) are OFF.

(2) The INST INDIRECT LIGHTS switch isON.

(3) The MASTER PANEL LIGHTS switch isOFF.

(4) The MASTER PANEL LIGHTS switch isON and the PILOT INST LIGHTS switch if OFF.

b. Master Warning Light (Red). MASTERWARNING light is located on each side of theglareshield (fig. 2-28). Any time a warning lightilluminates, the MASTER WARNING light willilluminate, and will remain illuminated until pressed. If anew fault condition occurs, the MASTER WARNINGlight will be reactivated, and the applicable annunciatorpanel warning light will illuminate.

c. Master Caution Light (Yellow). A MASTERCAUTION light is located on each side of the glareshield(fig. 2-28). Any time a caution light illuminates, theMASTER CAUTION will illuminate, and will remainilluminated until pressed. If a new fault conditionoccurs, the MASTER CAUTION light will be reactivated,and the applicable annunciator panel caution light willilluminate. Annunciator Panels

d. Pilot's Clock. One manual wind 8-day clock ismounted in the center of the pilot's control wheel (fig. 2-16).

e. Copilot's Clock. A digital quartz chronometer ismounted in the center of the copilot's control wheel (fig.2-16). The quartz chronometer is a five functionclock/timer that is controlled by three pushbuttonswitches located directly below the six digit liquid crystaldisplay.

(1) Mode selection. The MODE button ispressed to select the desired mode of operation. Themode annunciator is displayed above the modeidentifiers, and advances to indicate each of thefollowing modes:

• LC Local Time

• ZU Zulu or Greenwich Mean Time

• FT Trip or Flight Timer

• ET Elapsed Time

• DC Down counter with Alarm

(2) Local time mode (LC). Press the MODEbutton to advance the annunciator to LC. To set thehour, press the RST button once, then press and holdthe ADV button until the correct hour is displayed. Toset minutes, press the RST button again, then press andhold the ADV button

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Figure 2-28. Instrument Panel (Sheet 1 of 2)

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Figure 2-28. Instrument Panel (Sheet 2 of 2)

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TM 55-1510-219-10Table 2-6. Annunicator Panels

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TM 55-1510-219-10Table 2-6. Annunicator Panels (Continued)

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TM 55-1510-219-10Table 2-6. Annunicator Panels (Continued)

until the correct minute is displayed. Press the SETbutton once to display and hold selected time. Press theST-SP button to resume clock operation and/orsynchronize the display with a selected time standard.

(3) Zulu or Greenwich Mean Time mode(ZU). Press the MODE button to advance theannunciator to ZU and set time as for local time shownabove. Minutes and seconds do not need to be reset iflocal time is correctly set. Press the RST button todisplay minutes/seconds, then press again to activatethe complete display.

When changing time zones, the hour may bechanged as above. It is not necessary to change theminutes/seconds. Press the RST button twice to returnto the full time display.

(4) Trip/Flight timer mode (FT). Press theMODE button to advance the annunciator to FT. Pressand hold the ST-SP button and verify that the displayshows zero. The timer will activate at takeoff and stopat touchdown. To prevent an accidental reset of flighttime, the clock cannot manually reset during flight.

(5) Elapsed time mode (ET). Press theMODE button to advance the annunciator to ET.

Press the RST button to set the time display to zero.Press the ST-SP button one time. To stop the counting,press the ST-SP button a second time. Ending time willbe displayed until the RST button is pressed to clear thedisplay. The clock may be used in other modes and theelapsed time display will remain until cleared bypressing the RST button. If the timer is counting whenthe RST button is pressed, the display will reset to zeroand the count will begin again from zero.

(6) Downcounter mode (DC). Press theMODE button to advance the annunciator to DC. Pressthe SET button twice to reset the hour display to zero.Press and hold the ADV button until the desired hour isdisplayed. Press the SET button again, to reset theminute display to zero. Press and hold the ADV buttonuntil the desired minute is displayed. Press the SETbutton again, to reset the seconds display to zero.Press and hold the ADV button until the desired secondis displayed. Press the SET button again, to arm thecounter. Press the STSP button to begin countdown.When the countdown reaches zero, the display will flashfor approximately one minute and then reset. Thecountdown may also be reset at any time by pressingthe ST-SP button.

Section XII. SERVICING, PARKING, AND MOORING.

2-86. GENERALThe following paragraphs include the procedures

necessary to service the aircraft except lubrication. Thelubrication requirements of the aircraft are covered inthe aircraft maintenance manual. Tables 2-7, 2-8, 2-9and 2-10 are used for identification of fuel, oil, etc. usedto service the aircraft. The servicing instructionsprovide procedures and precautions necessary toservice the aircraft.

2-87. FUEL HANDLING PRECAUTIONS.

Table 2-2, Fuel Quantity Data, lists the quantityand capacity of fuel tanks in the aircraft. Service thefuel tanks after each flight to keep moisture out of thetanks and to keep the bladder type cells from drying out.Observe the following precautions:

WARNING

During warm weather open fuelcaps slowly to prevent beingsprayed with fuel.

WARNING

When aviation gasoline is used in aturbine engine, extreme caution shouldbe used when around the combustionchamber and exhaust area to avoid cutsor abrasions. The exhaust depositscontain lead oxide which will cause leadpoisoning.

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TM 55-1510-219-10

CAUTION

Proper procedures for handling JP-4 andJP-5 fuel cannot be over stressed. Clean,fresh fuel shall be used and the entranceof water into the fuel storage or aircraftfuel system must be kept to a minimum.

CAUTION

When conditions permit, the aircraft shallbe positioned so that the wind will carrythe fuel vapors away from all possiblesources of ignition. The fuel vehicle shallbe positioned to maintain a minimumdistance of 10 feet from any part of theaircraft, while maintaining a minimumdistance of 20 feet between the fuelingvehicle and the fuel filler point.

a. Shut off unnecessary electrical equipment on theaircraft, including radar and radar equipment. The masterswitch may be left on, to monitor fuel quantity gages, butshall not be moved during the fueling operation. Do notallow operation of any electrical tools, such as drills orbuffers, in or near the aircraft during fueling.

b. Keep fuel servicing nozzles free of snow, water,and mud at all times.

c. Carefully remove snow, frost, water, and ice fromthe aircraft fuel filler cap area before removing the fuelfiller cap (fig. 2-29). Remove only one aircraft tank fillercap at any one time, and replace each one immediatelyafter the servicing operation is completed.

d. Drain water from fuel tanks, filter cases, andpumps prior to first flight of the day. Preheat, whenrequired, to insure free fuel drainage.

e. Avoid dragging the fueling hose where it candamage the soft, flexible surface of the deicer boots.

f: Observe NO SMOKING precautions.

g. Prior to transferring the fuel, insure that the hoseis grounded to the aircraft.

h. Wash off spilled fuel immediately.

i. Handle the fuel hose and nozzle cautiously to avoiddamaging the wing skin.

j. Do not conduct fueling operations within 100 feetof energized airborne radar equipment or within 300 feet

of energized ground radar equipment installations.

k. Wear only nonsparking shoes near aircraft orfueling equipment, as shoes with nailed soles or metalheel plates can be a source of sparks.

2-88. FILLING FUEL TANKS.

WARNING

Prior to removing the fuel tank filler cap,the hose nozzle static ground wire shallbe attached to the grounding lugs thatare located adjacent to the filler opening.

Fill tanks as follows:

a. Attach bonding cables to aircraft.

b. Attach bonding cable from hose nozzle toground socket adjacent to fuel tank being filled.

CAUTION

Do not insert fuel nozzle completely intofuel cell due to possible damage tobottom of fuel cell.

c. Fill main tank before filling respective auxiliarytanks unless less than a full fuel load is desired.

d. Secure applicable fuel tank filler cap. Makesure latch tab on cap is pointed aft.

e. Disconnect bonding cables from aircraft.

2-89. DRAINING MOISTURE FROM FUEL SYSTEM.

To remove moisture and sediment from the fuelsystem, 12 fuel drains are installed (plus one for theferry system, when installed).

2-90. FUEL TYPES.

Approved fuel types are as follows:

a. Army Standard Fuels. Army standard fuel is JP-4.

b. Alternate Fuels. Army alternate fuels are JP-5and JP-8.

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Figure 2-29. Servicing Locations

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Table 2-7. Fuel, Lubricants, Specifications and Capacities

Table 2-8. Approved Fuels

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Table 2-8. Approved Fuels (Continued)

c. Emergency Fuel. Avgas is emergency fuel andsubject to 150 hour time limit.

2-91. USE OF FUELS.

Fuel is used as follows:

a. Fuel Mixture. Standard and alternate fuels maybe mixed in any ratio. Emergency fuels may be mixedin any ratio with standard and alternate fuels, however,use of the lowest octane rating available is suggested.Use of emergency fuel is subject to a 150 hour timelimit.

b. Use of Kerosene Fuels. The use of kerosenefuels (JP-5 type) in turbine engines dictates the need forobservance of special precautions. Both ground startsand air restarts at low temperature may be more difficultdue to low vapor pressure. Kerosene fuels having afreezing point of minus 40 degrees C (minus 40 degreesF) limit the maximum altitude of a mission to 28,000feet under standard day conditions.

c. Mixing of Fuels in Aircraft Tanks. Whenchanging from one type of authorized fuel to another, forexample JP-4 to JP-5, it is not necessary to drain theaircraft fuel system before adding the new fuel.

d. Fuel Specifications. Fuels having the sameNATO code number are interchangeable. Jet fuels

conforming to ASTM D-1655 specification may be usedwhen MIL-T-5624 fuels are not available. This usuallyoccurs during cross country flights where aircraft usingNATO F-44 (JP-5) are refueling with NATO F-40 (JP-4)or Commercial ASTM Type B fuels. Whenever thiscondition occurs, the engine operating characteristicsmay change in that lower operating temperature, sloweracceleration, lower engine speed, easier starting, andshorter range may be experienced. The reverse is truewhen changing from F-40 (JP-4) fuel to F-44 (JP-5) orCommercial ASTM Type A-1 fuels. Most commercialturbine engines will operate satisfactorily on eitherkerosene or JP-4 type fuel. The difference in specificgravity may possibly require fuel control adjustments; ifso, the recommendations of the manufacturers of theengine and airframe are to be followed.

2-92. SERVICING OIL SYSTEM.

An integral oil tank occupies the cavity formedbetween the accessory gearbox housing and thecompressor inlet case on the engine. The tank has acalibrated oil dipstick and an oil drain plug. Avoidspilling oil. Any oil spilled must be removedimmediately. Use a cloth moistened in solvent toremove oil. Overfilling may cause a discharge of oilthrough the accessory gearbox breather until asatisfactory level is reached. Service oil system asfollows:

Table 2-9. Standard, Alternate and Emergency.

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TM 55-1510-219-10

1. Open the access door on the uppercowling to gain access to the oilfiller cap and dipstick.

CAUTION

A cold oil check is unreliable. If possible,check oil within 10 minutes after engineshutdown. If over 10 minutes haveelapsed, motor the engine (starter only) for15-20 seconds, then recheck. If over 10hours have elapsed, start the engine andrun for 2 minutes, then recheck. Add oilas required. Do not overfill.

2. Remove oil filler cap.

3. Insert a clean funnel, with a screenincorporated, into the filler neck.

4. Replenish with oil to within 1 quartbelow MAX mark or the MAX COLDon dipstick (cold engine). Fill toMAX or MAX HOT (hot engine).

5. Check oil filler cap for damagedpreformed packing, generalcondition and locking.

CAUTION

Insure that oil filler cap is correctlyinstalled and securely locked to preventloss of oil and possible engine failure.

6. If oil level is over 2 quarts low,motor or run engine as required,and service as necessary.

7. Install and secure oil filler cap.

8. Check for any oil leaks.

2-93. SERVICING HYDRAULIC BRAKE SYSTEMRESERVOIR.

1. Gain access to brake hydraulicsystem reservoir.

2. Remove brake reservoir cap and fillreservoir to washer on dipstick withhydraulic fluid.

3. Install brake reservoir cap.

2-94. INFLATING TIRES.

Inflate tires as follows:

1. Inflate nose wheel tires to a pressurebetween 55 and.60 PSI.

2. Inflate main wheel tires to a pressurebetween 73 and 77 PSI.

2-95. SERVICING THE ELECTRIC TOILET.

The toilet should be serviced during routine groundmaintenance of the aircraft following any usage. It ismore efficient and convenient to remove, clean andrecharge the toilet tank on a regular basis than to wait untilthe tank is filled to capacity. Instructions for servicing areprovided on a decal applied to the front side of theremovable tank. Instructions are as follows:

a. Tank Removal.

1. Open front access to the toilet, asapplicable, to remove the toilet tank.

2. Depress the lock ring of the flush hosequick disconnect coupling located onthe right side at the front of the tanktop.

3. Drain any residue of flush fluid in thehose by partially disengaging the plugfrom the quick disconnect andmanipulating the hose to assistdrainage.

4. Remove the flush hose from the quickdisconnect and place hose in theretaining clip located on the undersideof the toilet mounting plate.

5. Install the cap attached to the quickdisconnect to seal the coupling.

6. Close the knife valve at the bottom ofthe toilet bowl by pushing the actuatorhandle until the valve is fully closed.

7. Press the two fasteners on each sideof the knife valve actuator to unlockthe tank.

8. Remove the tank by pulling therecessed carrying handle on the tanktop.

b. Tank Cleaning.

1. Dispose of tank contents by holdingthe tank upside-down over a sewer ortoilet and pull the knife valve actuatorhandle, opening the valve andallowing the tank to drain.

2. Rinse the tank by filling one-half fullwith water. Close the knife valve andshake vigorously. Drain tank as inprevious step.

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TM 55-1510-219-10

NOTE

Commercial detergents anddisinfectants can be included inthe rinse water if desired.However, do not include thesematerials in the tank precharge.

3. Rinse and drain the tank severaltimes to insure that the tank isthoroughly clean.

4. Wipe the exterior surfaces of thetank using a cloth moistened withclear water and disinfectant.

c. Tank Precharge.Caution

During freezing temperatures,toilet shall be serviced withantifreeze solution to preventdamage.

Charge the tank with a mixture of water andchemical according to chemical manufacturer'sspecification.

d. Tank Installation.

1. Install the tank by inserting theslides located on each side of theknife valve into the slide plateassembly on the bottom of thetoilet and slide tank into place.

2. Press the two fasteners to the firstdetent to secure the tank.

3. Remove the cap in the flush hosequick disconnect and connect thehose coupling to the quickdisconnect. Lock the disconnectlock ring.

4. Pull the knife valve actuator to fullyopen the valve.

5. Lift the toilet seat and shroudassembly from the top of the toiletand wipe with cloth moistened withclear water and disinfectant. Wipethe bowl and surrounding area.

6. Check flushing operation of thetoilet and check for leaks.

7. Close access to the toilet.

2-96. ANTI-ICING, DEICING AND DEFROSTINGPROTECTION.

The aircraft is protected in subfreezing weather byspraying the surfaces (to be covered with protective covers)with defrosting fluid. Spraying defrosting fluid on aircraftsurfaces before installing protective covers will permitprotective covers to be removed with a minimum of sticking.To prevent freezing rain and snow from blowing underprotective covers and diluting the fluid, insure that protectivecovers are fitted tightly. As a deicing measure, keepexposed aircraft surface wet with fluid for protection againstfrost.

NOTE

Do not apply anti-icing, deicing anddefrosting fluid to exposed aircraft surfacesif snow is expected. Melting snow willdilute the defrosting fluid and form a slushmixture which will freeze in place andbecome difficult to remove.

Use undiluted anti-icing, deicing, and defrosting fluid(MIL-A-8243) to treat aircraft surfaces for protection againstfreezing rain and frost. Spray aircraft surface sufficiently towet area, but without excessive drainage. A fine spray isrecommended to prevent waste. Use diluted, hot fluid toremove ice accumulations.

1. Remove frost or ice accumulations fromaircraft surfaces by spraying with dilutedanti-icing, deicing, and defrosting fluidmixed in accordance with Table 2-10.

2. Spray diluted, hot fluid in a solid stream(not over 15 gallons per minute).Thoroughly saturate aircraft surface andremove loose ice. Keep a sufficientquantity of diluted, hot fluid on aircraftsurface coated with ice to prevent liquidlayer from freezing. Diluted, hot fluidshould be sprayed at a high pressure,but not exceeding 300 PSI.

3. When facilities for heating are notavailable and it is deemed necessary toremove ice accumulations from aircraftsurfaces, undiluted defrosting fluid maybe used. Spray undiluted defrostingfluid at 15 minute intervals to assurecomplete coverage. Removal of iceaccumulations using undiluteddefrosting fluid in expensive and slow.

4. If tires are frozen to ground, useundiluted defrosting fluid to melt icearound tire. Move aircraft as soon astires are free. Recommended FluidDilution Chart

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2-97. APPLICATION OF EXTERNAL POWER.

CAUTIONBefore connecting the powercables from the external powersource to the aircraft, insure thatthe GPU is not touching theaircraft at any point. Due to thevoltage drop in the cables, thetwo ground systems will be ofdifferent potentials. Should theycome in contact while the GPU isoperating, arcing could occur.Turn off all external power whileconnecting the power cable to, orremoving it from the externalpower supply receptacle. Becertain that the polarity of theexternal power source is the sameas that of the aircraft before it isconnected. For GPU powerrequirements, consult theMaintenance Manual.

An external power source is often needed to supplythe electric current required to properly ground servicethe aircraft electrical equipment and to facilitate startingthe aircraft's engines. An external DC power receptacleis installed on the outboard side of the right enginenacelle. An external AC power receptacle is installed onthe outboard side of the left engine nacelle.

2-98. SERVICING OXYGEN SYSTEM.

The oxygen system furnishes breathing oxygen tothe pilot, copilot and first aid position. Figure 2-19shows the location of oxygen cylinder.

a. Oxygen System Safety Precautions.

WARNINGKeep fire and heat away fromoxygen equipment. Do not smokewhile working with or near oxygenequipment, and take care not togenerate sparks with carelesslyhandled tools when working onthe oxygen system.

(1) Keep oxygen regulators, cylinders, gages,valves, fittings, masks, and all other components of theoxygen system free of oil, grease, gasoline, and allother readily combustible substances. The utmost careshall be exercised in servicing, handling, and inspectingthe oxygen system.

(2) Do not allow foreign matter to enter oxygenlines.

(3) Never allow electrical equipment to come incontact with the oxygen cylinder.

(4) Never use oxygen from a cylinder withoutfirst reducing its pressure through a regulator.

b. Replenishing Oxygen System.1. Remove oxygen access door on outside of

aircraft (fig. 2-29).2. Remove protective cap on oxygen system

filler valve.3. Attach oxygen hose from oxygen servicing

unit to filler valve.

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Table 2-10. Recommended Fluid Dilution Chart

TM 55-1510-219-10

WARNINGIf the oxygen system pressure isbelow 200 PSI, do not attempt toservice system. Make an entry onDA Form 2408-13.

4. Insure that supply cylinder shutoff valveson the aircraft are open.

5. Fill system slowly to prevent overheating byadjusting recharging rate with pressureregulating valve on oxygen servicing unit.

6. Close pressure regulating valve on oxygenservicing unit when pressure gage onoxygen system indicates the pressureobtained using the Oxygen SystemServicing Pressure Chart (fig. 2-30).

NOTEFill oxygen system to 1800 PSI at70°F. For every degree above 70°,increase pressure 3.5 PSI to amaximum of 2000 PSI; and forevery degree F below 70°,decrease pressure 3.5 PSI.

7. Disconnect oxygen hose from oxygenservicing unit and filler valve.

8. Install protective cap on oxygen filler valve.9. Install oxygen access door.

2-99. GROUND HANDLING.Ground handling covers all the essential information

concerning movement and handling (fig. 2-31) of theaircraft while on the ground. The following paragraphsgive, in detail, the instructions and precautionsnecessary to accomplish ground handling functions.

a. General Ground Handling Procedure. Accidentsresulting in injury to personnel and damage toequipment can be avoided or minimized by closeobservance of existing safety standard and recognizedground handling procedures. Carelessness orinsufficient knowledge of the aircraft or equipment being

handled can be fatal. The applicable technical manualsand pertinent directives should be studied forfamiliarization with the aircraft, its components, and theground handling procedures applicable to it, beforeattempting to accomplish ground handling.

b. Ground Handling Safety Practice. Aircraftequipped with turboprop engines require additionalmaintenance safety practices. The following list ofsafety practices should be observed at all times toprevent possible injury to personnel and/or damaged ordestroyed aircraft:

(1) Keep intake air ducts free of loose articlessuch as rags, tools, etc.

(2) Stay clear of exhaust outlet areas.(3) During ground runup, make sure the brakes

are firmly set.(4) Keep area fore and aft of propellers clear of

maintenance equipment.(5) Do not operate engines with control

surfaces in the locked position.(6) Do not attempt towing or taxiing of the

aircraft with control surfaces in the locked position.(7) When high winds are present, do not unlock

the control surfaces until prepared to operate them.(8) Do not operate engines while towing

equipment is attached to the aircraft, or while the aircraftis tied down.

(9) Check the nose wheel position. Unless it isin the centered position, avoid operating the engines athigh power settings.

(10) Hold control surfaces in the neutral positionwhen the engines are being operated at high powersettings.

(11) When moving the aircraft, do not push onpropeller deicing boots. Damage to the heatingelements may result.

c. Moving Aircraft on Ground. Aircraft on theground shall be moved in accordance with the following:

(1) Taxiing. Taxiing shall be in accordance withChapter 8.

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Figure 2-30. Oxygen System Servicing Pressure

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Figure 2-31. Parking, Covers, Ground Handling, & Towing Equipment

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CAUTIONWhen the aircraft is being towed,a qualified person must be in thepilot's seat to maintain control byuse of the brakes. When towing,do not exceed nose gear turnlimits. Avoid short radius turns,and always keep the inside orpivot wheel turning during theoperation. Do not tow aircraftwith rudder locks installed, assevere damage to the nosesteering linkage can result. Whenmoving the aircraft backwards, donot apply the brakes abruptly.Tow the aircraft slowly, avoidingsudden stops, especially oversnowy, icy rough, soggy, ormuddy terrain. In arctic climates,the aircraft must be towed by themain gears, as an immensebreakaway load, resulting fromice, frozen tires, and stiffenedgrease in the wheel bearings maydamage the nose gear.

CAUTIONDo not tow or taxi aircraft withdeflated shock struts.

(2) Towing. Towing lugs are provided on theupper torque knee fitting of the nose strut. When it isnecessary to tow the aircraft with a vehicle, use thevehicle tow bar. In the event towing lines arenecessary, use towing lugs on the main landing gear.Use towing lines long enough to clear nose and/or tail byat least 15 feet. This length is required to prevent theaircraft from overrunning the towing vehicle or foulingthe nose gear.

d. Ground Handling Under Extreme WeatherConditions. Extreme weather conditions necessitateparticular care in ground handling of the aircraft. In hot,dry, sandy, desert conditions, special attention must bedevoted to finding a firmly packed parking and towingarea. If such areas are not available, steel mats or anequivalent solid base must be provided for thesepurposes. In wet, swampy areas, care must be taken toavoid bogging down the aircraft. Under cold, icy, arcticconditions, additional mooring is required, and addedprecautions must be taken to avoid skidding duringtowing operations. The particular problems to beencountered under adverse weather conditions and thespecial methods designed to avoid damage to the

aircraft are covered by the various phases of the groundhandling procedures included in this section of generalground handling instructions. (Refer to TM 55-1500-204-25/1.)

2-100. INSTALLATION OF PROTECTIVE COVERS.The crew will insure that the aircraft protective

covers are installed.

2-101. MOORING.The aircraft is moored to insure its immovability,

protection, and security under various weatherconditions. The following paragraphs give, in detail, theinstructions for proper mooring of the aircraft.

a. Mooring Provisions. Mooring points (fig. 2-32)are provided beneath the wings and tail. Additionalmooring cables may be attached to each landing gear.General mooring equipment and procedures necessaryto moor the aircraft, in addition to the following, aregiven in TM 55-1500-204-25/1.

(1) Use mooring cables of 1/4 inch diameteraircraft cable and clamp (clip-wire rope), chain or rope3/8 inch diameter or larger. Length of the cable or ropewill be dependent upon existing circumstances. Allowsufficient slack in ropes, chains, or cable to compensatefor tightening action due to moisture absorption of ropeor thermal contraction of cable or chain. Do not use slipknots. Use bowline knots to secure aircraft to mooringstakes.

(2) Chock the wheels.b. Mooring Procedures for High Winds. Structural

damage can occur from high velocity winds; therefore, ifat all possible, the aircraft should be moved to a safeweather area when winds above 75 knots are expected.If aircraft must be secured use the following steps:

1. After aircraft is properly located, place nosewheel in centered position. Head aircraftinto the wind, or as nearly so as is possiblewithin limits determined by locations of fixedmooring rings. When necessary, a 45degree variation of direction is considered tobe satisfactory. Locate each aircraft atslightly more than wing span distance fromall other aircraft. Position nose mooringpoint approximately 3 to 5 feet downwindfrom ground mooring anchors.

2. Deflate nose wheel shock strut to within 3/4inch of its fully deflated position.

3. Fill all fuel tanks to capacity, if time permits.4. Place wheel chocks fore and aft of main gear

wheels and nose wheel. Tie each pair ofchocks together with rope

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Figure 2-32. Mooring the Aircraft

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or join together with wooden cleats nailed tochocks on either side of wheels. Tie ice gripchocks together with rope. Use sandbags inlieu of chocks when aircraft is moored onsteel mats. Set parking brake as applicable.

5. Accomplish aircraft tiedown by utilizingmooring points shown in figure 2-32. Maketiedown with 1/4 inch aircraft cable, usingtwo wire rope clips or bolts, and a chairtested for a 3000 pound pull. Attachtiedowns so as to remove all slack. (Use a3/4-inch or larger manila rope if cable orchain tiedown is not available.) If rope isused for tiedown, use anti-slip knots, suchas bowline knot, rather than slip knots. Inthe event tiedown rings are not available onhard surfaced areas, move aircraft to anarea where portable tiedowns can be used.Locate anchor rods at point shown in figure2-31. When anchor kits are not available,use metal stakes or deadman type anchors,providing they can successfully sustain aminimum pull of 3000 pounds.

6. In event nose position tiedown is consideredto be of doubtful security due to existing soilcondition, drive additional anchor rods atnose tiedown position. Place padded workstand or other suitable support under the aftfuselage tiedown position and secure.

7. Place control surfaces in neutral position.Place wing flaps in up position.

8. The requirements for dust excluders,protective covers, and taping of openingswill be left to the discretion of theresponsible maintenance officer or the pilotof the transient aircraft (fig. 2-31).

9. Secure propellers to prevent windmilling.10. Disconnect battery.11. During typhoon or hurricane wind conditions,

mooring security can be further increasedby placing sandbags along the wings tobreak up the aerodynamic flow of air overthe wing, thereby reducing the lift beingapplied against the mooring by the wind.The storm appears to pass two times, eachtime with a different wind direction. This willnecessitate turning the aircraft after the firstpassing.

12. After high winds, inspect aircraft for visiblesigns of structural damage and for evidence

of damage from flying objects. Servicenose shock strut and reconnect battery.

2-102. PARKING.Parking is defined as the normal condition under

which the aircraft will be secured while on the ground.This condition may vary from the temporary expedientof setting the parking brake and chocking the wheels tothe more elaborate mooring procedures described underMooring. The proper steps for securing the aircraft mustbe based on the time the aircraft will be left unattended,the aircraft weights, the expected wind direction andvelocity, and the anticipated availability of ground andair crews for mooring and/or evacuation. Whenpractical head the aircraft into the wind, especially ifstrong winds are forecast or if it will be necessary toleave the aircraft overnight. Set the parking brake andchock the wheels securely. Following engine shutdown,position and engage the control locks.

NOTECowlings and loose equipmentwill be suitably secured at alltimes when left in an unattendedcondition.

a. The parking brake system for the aircraftincorporates two lever-type valves, one for each wheelbrake. Both valves are closed simultaneously by pullingout the parking brake handle. Operate the parkingbrake as follows:

1. Depress both brakes.2. Pull parking brake handle out. This will

cause the parking brake valves to lock thehydraulic fluid under pressure in the parkingbrake system, thereby retaining brakingaction.

3. Release brake pedals.

CAUTIONDo not set parking brakes whenthe brakes are hot during freezingambient temperatures. Allowbrakes to cool before settingparking brakes.

4. To release the parking brakes push in on theparking brake handle.

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b. The control lock (fig. 2-17) holds the engine andpropeller control levers in a secure position, and theelevator, rudder, and aileron in neutral position. Installthe control locks as follows:

1. With engine and propeller control levers insecure position, slide lock onto controlpedestal to prevent operation of levers.

2. Install elevator and aileron lockpin verticallythrough pilot's control column to lock controlwheel.

3. Install rudder lock pin through flapper doorforward to pilot's seat, making sure rudder isin neutral position.

4. Reverse steps 1 through 3 above to removecontrol lock. Store control lock.

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CHAPTER 3AVIONICS

Section I. GENERAL

3-1. INTRODUCTION.Except for mission avionics, this chapter covers all

avionics equipment installed in the RC-12D aircraft. Itprovides a brief description of equipment covered, thetechnical characteristics and locations. It coverssystems and controls and provides the propertechniques and procedures to be employed whenoperating the equipment. For more detailed operationalinformation consult the vendor manuals that accompanythe aircraft loose tools.

3-2. AVIONICS EQUIPMENT CONFIGURATION.The aircraft avionics covered is comprised of three

groups of electronic equipment. The Communicationgroup consists of the Interphone, UHF command, VHFcommand and HF command systems. The Navigationgroup provides the pilot and copilot with theinstrumentation required to establish and maintain anaccurate flight course and position, and to make anapproach on instruments under InstrumentMeteorological Conditions (IMC). The Navigation groupincludes equipment for determining altitude, attitude,position, destination, range and bearing, headingreference, groundspeed, and drift angle. TheTransponder and Radar group includes an identification,position and emergency tracking system, and a radarsystem to locate potentially dangerous weather areasand a radar system to differentiate between friendly andunfriendly search radar.

NOTEAll avionics equipment require a3-minute warmup period. Theweather radar has an automatictime delay of 3 to 4 minutes.

3-3. POWER SOURCE.a. DC Power. DC power for the avionics equipment

is provided by four sources: the aircraft battery, left andright generators, and external power. Power is routedthrough a 50-ampere circuit breaker to the avionicspower relay which is controlled by the AVIONICSMASTER POWER SWITCH (fig. 2-18) on the overhead

control panel. Individual system circuit breakers and theassociated avionics busses are shown in fig. 2-22.With the switch in the ON position, the avionics powerrelay is de-energized and power is applied through boththe AVIONICS MASTER POWER #1 and #2 circuitbreakers to the individual avionics circuit breakers onthe overhead circuit breaker panel (fig. 2-26). In the off(aft) position, the relay is energized and power isremoved from avionics equipment. When externalpower is applied to the aircraft, the avionics power relayis normally energized, removing power from theavionics equipment. To apply external power to theavionics equipment, move the AVIONICS MASTERPOWER switch to the EXT PWR position. This will de-energize the avionics power relay and allow power to beapplied to avionics equipment.

b. AC Power. AC power for the avionics equipmentis provided by two inverters. The inverters supply 115-volt and 26-volt single-phase AC power when operatedby the INVERTER #1 or #2 switches (fig. 2-18). Eitherinverter is capable of powering all avionics equipmentrequiring AC power. AC power from the inverters isrouted through fuses in the nose avionics compartment.

3-4. MICROPHONES, SWITCHES AND JACKS.Boom, and oxygen mask microphones can be utilized inthe aircraft.

a. Microphone Switches. The pilot and copilot areprovided with individual MIC control switches, placardedMIC/INTPH/XMIT, attached to respective controlwheels. A foot-actuated mic switch is also positioned onthe floorboards forward of each pilot's seat.

b. Controls and Functions. Microphone switchesand jack functions are as follows:

(1) Control wheel switches (fig. 2-16).

CONTROL FUNCTIONMIC/INTPH Keys selected facility.XMIT switchMIC (Not depressed) Microphone is

disconnected.

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INTPH (depressed to first detent) Keysinterphone facility, disregardsposition transmitter selectorswitch.

XMIT (depressed full down) Keys facil-ity selected by transmitter selectswitch.

(2) Floorboard switches.CONTROL FUNCTIONMIC foot switch Controls connection of selected

mic to audio system.Held depressed Connects selected mic to audio

system.

Released Disconnects selected mic fromaudio system.

(3) Subpanel jack selector switches.CONTROL FUNCTIONMIC HEADSET/ Selects microphone to connectOXYGEN to audio system switchMASKMIC HEADSET Connects headset microphone to

audio system.OXYGEN Connects microphone of oxygenMASK mask to audio system.

Section II. COMMUNICATIONS

3-5. DESCRIPTION.The communications equipment group is comprised

of an interphone system connected to individual AudioControl Panels for the pilot and copilot which interfacewith VHF, UHF and HF communication units.

3-6. AUDIO CONTROL PANELS (FIG. 3-1).a. Description. Separate but identical audio control

panels serve the pilot and copilot. The controls andswitches of each panel provide the user with a means toselect desired reception and transmission sources, andalso a means to control the volume of audio signals sentand received for interphone, communication andnavigation systems. The user selects between the UHF,VHF and HF transceivers. Audio control panels areprotected by respective 2-ampere AUDIO PILOT andAUDIO COPILOT circuit breakers positioned on theoverhead circuit breaker panel (fig. 2-26).

b. Controls and Indicators.(1) Audio controls - Comm and Nav Radios (fig.

3-1).CONTROL/ FUNCTIONINDICATORMaster VOL Controls sidetone volume tocontrol headset. Also serves as final

volume adjustment for receivedaudio from any source beforeadmission to headset.

Comm radios Each is combination rotary con-audio controls trol and "ON-OFF" push-pull

switch, permitting both receiver selection and volume adjustment.

1 ON connects user's headset toaudio from VHF-AM transceiverNo. 1.

2 ON connects user's headset toaudio from VHF/AM/FM trans-ceiver.

3 ON connects user's headset toaudio from command UHFtransceiver in use.

4 ON connects user's headset toaudio from HF or VOW trans-ceiver in use.

5 ON connects user's headset toaudio from backup VOW trans-ceiver.

NAV radios Combination volume controlaudio monitor and "ON-OFF" switches forcontrols NAV receivers.NAV-A ON connects user's headset to

audio from VOR-1, VOR-2 orset in use.

NAV-B ON connects user's headset toaudio from TACAN or ADF setin use.

(2) Microphone switches.CONTROL FUNCTIONMic impedance Two-position thumb-actuatedselect switch (5 switch. Enables selection of in-Ohm/150 Ohm) terface circuit with best imped-

ence match to microphone used.

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TM 55-1510-219-10

Microphone Controls activation of micro-select switches phones.HOT MIC Admits speech to interphone

system without need to key se-lected microphone.

NORM Blocks speech from interphonesystem unless selected micro-phone is keyed.

ICS OFF Deactivates microphones(3) Transmitter select switch (fig. 3-1).

CONTROL FUNCTIONTransmit Connects microphone and head-interphone set to selected radio transmitterselector switch or interphone line routing re-

ceived audio to headset. Bypass-es control of respective receiveraudio switch.

PVT Position not used.ICS Activates pilot-to-copilot inter-

com.1 Permits audio reception from

VHF-AM No. 1 transceiver.Routes key and mic signals toVHF-AM No. 1 transceiver.

2 Permits audio reception fromVHF/AM/FM transceiver.Routes key and mic signals toVHF/AM/FM transceiver.

3 Permits audio reception fromUHF transceiver. Routes keyand mic signals to UHF trans-ceiver.

4 Permits audio reception fromHF or VOW transceiver. Routeskey or mic signals to transceiver.

5 Permits audio reception frombackup VOW. Routes key andmic signals to transceiver.

(4) CIPHER light - not used.c. Normal Operation.

(1) Turn-on. Both Audio Control Panels areactivated when electrical power is applied to aircraft.

NOTEIt is presumed the AVIONICSMASTER POWER switch is on,

and that normally used avioniccircuit breakers remaindepressed. The circuit breakersof routinely used avionic systemsare normally left depressed.

(2) Receive.1. Receiver audio switches (Audio panel) -

As required.2. Master VOL control (Audio panel) - As

required, in combination with volumecontrol of system being utilized.

NOTEAudio select switches and volumecontrols are routinely left inpositions of normal use.

3. Move each receiver audio switch ONthen OFF, separately, to verify audiopresence in headphones for eachsystem. Adjust volume.

4. Mic select switch - As desired.(3) Transmit.

1. Transmitter-interphone selector (audiopanel) - Set for transceiver desired.

2. MIC HEADSET/OXYGEN MASK switch(instrument panel) -As desired.

3. MIC/INTPH/XMIT switch (control wheel)- XMIT.

4. MIC switch (headset/oxygenmask/floorboards) -Depress to transmit.

(4) Intercommunication.1. Transmitter-interphone selector (audio

panel) -ICS.2. MIC HEADSET/OXYGEN MASK switch

(instrument panel) -As desired.3. Mic select switch (audio panel - As

desired (NORM/HOT MIC).a. If HOT MIC selected - Talk when

ready.b. If NORM selected - Depress MIC

switch and transmit.c. Master VOL control (selected

transceiver) -Set for comfort.

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Figure 3-1. Audio Control Panel (C-499) (Typical pilot, copilot)

d. Emergency Operation. Not applicable.e. Shutdown.

1. AVIONICS MASTER POWER switch(overhead panel) -OFF.

2. Leave avionic controls and circuitbreakers positioned for normal operation.

3. Aircraft DC power - OFF.

3-7. RADIO CONTROL PANEL (FIG. 3-2).a. Description. The radio control panel, located on

the pedestal extension allows the pilot or copilot toselectively monitor audio signals from the HighFrequency (HF), Marker Beacon (MKR BCN), or ADFsystems. It also has controls for the selection of ADFvoice or range filters. Controls and functions are asfollows:

b. Controls and Indicators.CONTROL FUNCTIONHF VOL control Adjusts volume of high-

frequency radio signals received.

MKR BCN HI/ Selects sensitivity of marker bea-LO switch con receiver.MKR BCN VOL Adjusts volume of marker bea-

control con radio signals received.(2) Automatic Direction Finding (ADF).

FILTER V/OFF Selects filter to block voiceswitch transmissions from ADF groundstation.FILTER R/OFF Selects filter to block rangeswitch transmissions from ADF ground

station.

3-8. UHF COMMAND SET (AN/ARC-164).a. Description. The UHF command set is a line-of-

sight radio transceiver which provides transmission andreception of amplitude modulated (AM) signals in theultra high frequency range of 225.000 to 399.975 MHzfor a distance range of approximately 50 miles.Channel selection is spaced at 0.025 MHz. A separatereceiver is incorporated to provide monitoring capabilityfor the UHF guard frequency (243.0 MHz). UHF audiooutput is applied to the audio panel where it is routed tothe headsets.

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TM 55-1510-219-10

Figure 3-2. Radio Control Panel

NOTEThe PRESET channel selector andmanual frequency selectors areinoperative when the modeselector is set to GUARD position.The receiver-transmitter will beset to the emergency frequencyonly.

The transmit and receive sections of the UHF unitoperate independently but share the same power supplyand frequency control circuits.

Complete provisions only are installed for voicesecurity device KY-28 or KY-58 to locate on thepedestal extension near the radio set. The UHFcommand set is protected by the 7 1/2 ampere UHFcircuit breaker on the overhead circuit breaker panel(fig. 2-26). Figure 3-3 illustrates the UHF commandset. The associated blade type antenna is shown infigure 2-1.

b. Controls and Indicators.(1) UHF control panel fig. 3-3.

CONTROL FUNCTIONManual Selects hundreds digit of fre-frequency quency (either 2 or 3) in MHz.selector

(hundreds)Manual Selects tens digit of frequency (Ofrequency through 9) in MHz.selector (tens)Manual Selects units digit of frequencyfrequency (O through 9) in MHz.selector (units)Manual Selects tenths digit of frequencyfrequency (O through 9) in MHz.selector (tenths)Manual Selects hundredths and thou-frequency sandths digits of frequency (00,selector 25, 50, or 75) in MHz.(hundredths andthousandths)Preset channel Selects one of 20 preset channelselector frequencies.Mode selector Selects operating mode and

method of frequency selection.MANUAL Enables the manual selection of

any one of 7,000 frequencies,PRESET Enables selection of any one of

20 preset channels. preset chan-nel selector switch.

GUARD Selection automatically tunes themain receiver and transmitter to

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Figure 3-3. UHF Control panel (AN/ARC- 164)

the guard frequency and theguard receiver is disabled.

SQUELCH Turns main receiver squelch onswitch or off.VOL control Adjusts volume.TONE When pressed, transmits a 1,020pushbutton Hz tone on the selected frequen-

cy.Function selector Selects operating function.OFF Turns set off.MAIN Selects normal transmission with

reception on main receiver.BOTH Selects normal transmissions

with reception on both the mainreceiver and the guard frequencyreceiver.

ADF Activates ADF or homing sys-tem (if installed) and main re-ceiver.

c. Normal Operation.(1) Turn on.

1. Insure aircraft power is on.

NOTE

It is presumed the AVIONICSMASTER POWER switch is on,and that normally used avioniccircuit breakers remaindepressed.

2. AVIONICS MASTER POWER switch(overhead panel) - ON.

3. Function switch (UHF panel) MAIN orBOTH position, as required.

NOTEIf function selector is at MAINsetting, only the normal UHFcommunications will be received.If selector is at BOTH position,emergency communications onthe guard channel and normalUHF communications will both bereceived.

(2) Receive.1. UHF audio monitor switch (#3, audio

panel) - ON or transmit/interphoneselector (audio panel) - 3 position.

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TM 55-1510-219-10

2. VOL control (UHF panel) - Mid position.(3) To use preset frequency (UHF panel).

1. Mode selector - PRESET position.2. PRESET channel selector - Rotate to

channel desired.(4) To use non-preset frequency (UHF panel).

1. Mode selector - MANUAL position.2. Manual frequency selectors (5) - Rotate

each knob to set desired frequencydigits.

NOTEThe PRESET channel selector andmanual frequency selectors areinoperative when the modeselector is set to GUARD position.

(5) Volume - Adjust.

NOTETo adjust volume when audio isnot being received, turn squelchswitch OFF, adjust volume forcomfortable noise level, then turnsquelch switch ON.

(6) Squelch - As desired.(7) Transmit.

1. Transmitter-interphone selector (audiopanel) - 3 position.

2. UHF panel - Set required frequencyusing either preset channel control ormanual frequency select controls.

3. MIC HEADSET/OXYGEN MASK switch(instrument panel) -As desired.

4. MIC switch - Depress to transmit.

(8) Shutdown.1. Function selector (UHF panel) - OFF.

d. UHF Command and Voice Security Operation(KY-28).

NOTEDisregard operating proceduresinvolving the voice security(CIPHONY) control-indicator, ifunit is not installed.

(1) Turn-on.1. POWER ON switch (Voice Security

panel, fig 3-6) -ON.2. Function switch (UHF panel) - BOTH.

(2) Receive (UHF panel).1. Mode selector - As required.2. Transmitter-interphone selector (audio

panel, fig. 3-1) - #3 position. or Audiomonitor control #3 - ON.

3. Set required frequency using presetchannel control or manual frequencyselector.

4. Adjust volume.

NOTETo adjust volume when radio isnot being received, turnSQUELCH switch OFF, adjustvolume for comfortable noiselevel, then turn squelch disableswitch ON.

5. Squelch - As required.(3) Transmit (PLAIN).

1. Transmit-interphone selector (audiopanel) - No. 3 position.

2. PLAIN/CIPHER switch (Voice Securitypanel) - PLAIN.

3. Microphone switch - Press.(4) Transmit (CIPHER).

1. Transmit-interphone selector (audiopanel) - No. 3 position.

2. PLAIN/CIPHER switch (Voice securitypanel) -CIPHER. (CIPHER indicator willbe on as long as switch is in CIPHERposition.)

3. RE-X/REG switch (Voice security panel)- REG.

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4. Microphone switch - Press momentarily(interrupted tone from voice securityunit should no longer be heard).

NOTENo traffic will be passed if theinterrupted tone is still heard afterpressing and releasing themicrophone switch.

5. Microphone switch - Press (do not talk).Wait until beep is heard then speak intomicrophone.

(5) Shutdown.1. Function switch (UHF panel)-OFF.2. POWER ON switch (Voice security

panel) - OFF.e. UHF Command Set - Emergency Operation.

NOTETransmission on emergencyfrequencies (guard channels) isrestricted to emergencies only.An emergency frequency of 121.500 MHz is also available on theVHF command radio set.

1. Transmit/interphone selector (audiopanel) - No. 3 position.

2. Mode selector switch (UHF panel) -GUARD.

3. Microphone switch - Press.

3-9. BACKUP VOW (AN/ARC-164).A radio set identical in type and performance to the

UHF command set (fig. 3-3) is installed on the pedestalextension (fig. 2-10) to serve as backup VOW (VoiceOrder Wire). This set provides the pilot and copilot withsecure 2-way voice communications. Completeprovisions only are provided for a KY-58 voice securitydevice. The backup VOW set is protected by a 7 1/2ampere BU VOW circuit breaker on the overhead circuitbreaker panel (fig. 2-26). The backup VOW andTransponder share an antenna which is mounted on theaircraft belly (fig. 2-1).

3-10. VHF-AM COMMUNICATIONS (VHF-20B).a. Description. VHF-AM communications provide

transmission and reception of amplitude modulatedsignals in the very high frequency range of 116.000 to151.975 MHz for a range of approximately 50 miles,varying with altitude. A dual head control panel (COMM1, fig. 3-4) is mounted on the pedestal (fig. 2-10)accessible to both the pilot and copilot. The panelprovides two sets of control indicators, frequencyindicators, frequency select knobs, a single volumecontrol, and a single selector switch for quick frequencychanging. Transmission audio is routed by pilot andcopilot #1 transmitter select switches. Received audio isrouted by pilot and copilot # receiver audio switches tothe respective headsets. The VHF radio is protected byone 10-ampere VHF circuit breaker on the overheadcontrol circuit breaker panel (fig. 2-26). The associatedantenna is shown in figure 2-1.

b. Controls and Indicators.(1) VHF-AM control panel (fig. 3-4).

CONTROL FUNCTIONLeft control Indicates set operating frequencyFrequency (control TRANS switch left posi-indicator tion.)Frequency Selects desired set operating fre-selectors quency (control TRANS switch

left position.)VOL-OFF Adjusts volume of received au-control dio, turns set ON or OFF.CONTROL Illuminates, if control TRANSindicator switch in left position.TRANS switch Selects right or left control head

to control operating frequency ofset.

Frequency Indicates set operating frequencyindicator (control TRANS switch right po-

sition.)Frequency Selects desired set operating fre-selectors quency (control TRANS switch

right position.)CONTROL Illuminates, if control TRANSindicator switch at right position.COMM TEST Overrides automatic squelch cir-switch cuit.

c. VHF-AM Set - Normal Operation.(1) Turn-on procedure. VOL control - Turn

clockwise.

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Figure 3-4. VHF-AM Control Panel (VHF-20B)(2) Receiver operating procedure.

1. Transmitter-interphone selector (audiopanel, fig. 3-1) - #1 position. or Audiomonitor control #1 - ON.

2. Frequency selector - Set desiredfrequency.

3. VOL control - As required.(3) Transmitter operating procedure.

1. Transmitter-interphone selector (audiopanel, fig. 3-1) - #1 position.

2. Microphone switch - Press.(4) Shutdown procedure. VOL control -

Counter-clockwise (OFF).d. VHF-AM Set - Emergency Operation.

NOTETransmissions on frequency(121.500 MHz) are restricted toemergencies only. Emergencyfrequency 243.000 MHz (guardchannel) is also available on theUHF command radio.

1. Transmitter-interphone selector (audio panel, fig. 3-1) - #1 position.

2. Frequency selector (VHF panel) - 121.500 MHz (emergency frequency).

3. Microphone switch - Press.

3-11. VHF AM/FM COMMAND SET (AN/ARC-186).a. Description. The VHF AM/FM Command Set

provides for normal and secure 2-way voicecommunication: AM in the very high frequency range of116.000 to 151.975 MHz and FM in the 30. 000 to87.975 MHz band. Twenty channels may be preset.Audio signals are applied through pilot and copilottransmitter select switches No. 2 position and throughthe pilot and copilot No. 2 receiver audio switches totheir respective headsets. Complete provisions only areinstalled for voice security device KY-28 or KY-58.Circuits are protected by a 10-ampere VHF AM/FMcircuit breaker on the overhead circuit breaker panel(fig. 2-26). Figure 3-5 illustrates the VHF AM/FMcontrol panel. The associated antenna is shown infigure 2-1.

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Figure 3-5. VHF AM/FM Control Panel (AN/ARC- 186)

b. Controls and Indicators.(1) VHF AM/FM control panel (fig. 3-5).

CONTROL FUNCTION10 MHz selector Selects receiver-transmitter fre-

quency in increments of 10 MHzfrom 30 to 150 MHz. Clockwiserotation increases frequency.

10 MHz Indicates manually selected re-indicator ceiver-transmitter frequency in

10 MHz increments from 30 to150 MHz.

1.0 MHz selector Selects receiver-transmitter fre-quency in 1.0 MHz increments.Clockwise rotation increases fre-quency.

1.0 MHz Indicates manually selected re-indicator ceiver-transmitter frequency in

1.0 MHz increments.0.1 MHz Indicates manually selected re-indicator ceiver-transmitter frequency in

0.1 MHz. increments.0.1 MHz selector Selects receiver-transmitter fre-

quency in 0.1 Mhz increments.Clockwise rotation increases fre-quency.

0.025 MHz Indicates manually selected re-indicator ceiver-transmitter frequency in

0.025 MHz increments.0.025 MHz Selects receiver-transmitter fre-selector quency in 0.025 MHz incre-

ments. Clockwise rotation in-creases frequency.

CHAN indicator Selects preset channel from 1 to20. Clockwise rotation increasesnumber selected.

LOCKOUT FM/ Screwdriver adjustable three-AM switch position switch. Warning tone

announces lockout.Center Selects AM or FM band.AM Disables AM band.FM Disables FM band.FM SQUELCH Screwdriver adjustable potenti-control ometer. Squelch fully overdriven

at full counter-clockwise posi-tion. Clockwise rotation in-creases input signal required toopen squelch.

WB/NB MEM Three-position switch.LOAD switch

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TM 55-1510-219-10

NB Limits selectivity to narrow-band intermediate frequency.

WB Limits selectivity to wide-bandintermediate frequency of FMband.

MEM LOAD Momentary switch. If pressed,loads manually selected frequen-cy in preset channel memory

AM SQUELCH Screwdriver adjustable potenti-control ometer. Squelch fully overridden

at full counter-clockwise posi-tion. Clockwise rotation in-creases input signal required toopen squelch.

Mode select Three-position rotary switch.switch

OFF Shuts off receiver-transmitter.TR Selects transmit/receive modes.DF Not used.SQ/DIS/TONE Three-position switch.

select switchCenter Selects squelch function,SQ/DIS Shuts off squelch function.TONE Transmits tone of approx. 1000

Hz.Frequency Four-position switch.

controlemergency/selectswitch

PRE Enables preset channel selection.MAN Enables manual frequency selec-

tion.EMER AM or Selects a prestored guard chan-FM nel.VOL control Clockwise rotation increases

volume.c. Normal Operation.

(1) Turn-on.1. Mode select switch (VHF AM/FM panel) -

TR.

(2) Receive.1. Frequency control emergency select

switch (fig. 3-5) - MAN or PRE, asdesired.

2. Transmitter-interphone selector (audiopanel, fig. 3-1) - #2 position. or Audiomonitor control #2 - ON.

3. Manual frequency/preset channelselectors - Set desired frequency.

4. VOL control - As required.(3) Transmit.

1. Transmit/interphone select switch (audiopanel) -Position 2.

2. Microphone switch - Press.(4) Shutdown.

1. Mode select switch (fig. 3-5) - OFF.d. VHF AM/FM - Emergency Operation.

(1) Emergency AM Mode.1. Transmit/interphone select switch (audio

panel) -Position 2.2. Mode select switch - TR.3. PRE-MAN-FM/AM EMER switch -

EMER AM.4. Manual frequency/preset channel

selector - Set emergency frequency.5. Microphone switch - Press.

(2) Emergency FM Mode.1. Transmit-interphone selector (audio

panel, fig. 3-1) - #2 position.2. Mode select switch - TR or DF, as

desired.3. PRE-MAN-FM/AM EMER switch -

EMER FM.4. Manual frequency/preset channel

selector - Set emergency frequency.5. Microphone switch - Press.6. Shutdown Mode select switch - OFF.

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3-12. VOICE SECURITY SYSTEM TSEC/KY-28(PROVISIONS ONLY).

NOTEVoice security system TSEC/KY-58 may be installed in lieu ofvoice security system TSEC/KY-28. Complete provisions areprovided to install either the KY-28 or the KY-58 voice securitysystem on the pedestal extension(fig. 2-10).

a. Description. The KY-28 voice security systemprovides secure (ciphered) two-way voicecommunications for the pilot and copilot in conjunctionwith the UHF and VHF/AM/FM command sets, and thebackup VOW set. System circuits are protected by theVHF, VHF/AM/FM, RADIO RELAY, and BU VOW circuitbreakers on the overhead circuit breaker panel (fig. 2-26). Figure 3-6 illustrates the KY-28 voice security(CIPHONY) control indicator.

b. Controls and Indicators.(1) Voice security control/indicator. (fig. 3-6).

CONTROL FUNCTIONINDICATORPOWER ON Turns set on or off.switch

NOTEThe POWER ON switch must be inON position for FM liaison orsecure mission operations ineither the plain or cipher mode.

POWER ON Illuminates when POWER ONindicator switch is placed in ON (up) position.PLAIN indicator Illuminates when PLAIN/

CIPHER switch is in PLAIN po-sition.

PLAIN/CIPHERswitchPLAIN Enables unciphered communica-

tions on FM liaison set.CIPHER Enables ciphered communica-

tions on FM liaison set.RE-X/REGswitchRE-X Enables re-transmission of ci-

phered communications at a dis-tant location.

REG Enables normal cipher or plaincommunications.

ZEROIZE switch Normally OFF. Place in ON po-sition during emergency situa-tions to neutralize and make in-operative the associated cipherequipment.

CIPHER Illuminates when PLAIN/indicator CIPHER switch is in CIPHER

position.c. VHF/AM/FM Set and Voice Security Operation.(1) Turn-on.

1. POWER ON switch (fig. 3-6) - ON.

NOTEThe POWER ON switch must be inON position, regardless of themode of the operation, wheneverthe voice security (CIPHONY) KY-28 is installed in the aircraft.

(2) Receive.1. SQUELCH control (VHF/AM/ FM

panel) - As required.2. Transmitter-interphone selector (audio

panel, fig. 3-1) - #2 position. or Audiomonitor control #2 - ON.

3. Mode selector (VHF/AM/FM panel) -TR.

4. Frequency selectors (VHF/AM/FMpanel) - As required. PLAIN/CIPHERswitch (voice security panel) - Asrequired.

(3) Transmit (PLAIN).1. Transmit/interphone selector (audio

panel) - No. 2 position.2. PLAIN/CIPHER switch (Voice security

panel) - PLAIN.3. Microphone switch - Press.

(4) Transmit (CIPHER).1. Transmit/interphone selector (audio

panel) - No. 2 position.2. PLAIN/CIPHER switch (Voice security

panel) -CIPHER. Indicator will be onwhile switch is in CIPHER position.)

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Figure 3-6. Voice Security Control Indicator (C-8157/ARC) (KY-28)

3. RE-X/REG switch (Voice security panel)- As required. (Set RE-X position only ifdistant station is using re-transmittingequipment.)

4. Microphone switch - Press momentarily(interrupted tone from voice securityunit should no longer be heard).

NOTENo traffic will be passed if theinterrupted tone is still heard afterpressing and releasing themicrophone switch.

5. Microphone switch- Press (do not talk).Wait until beep is heard, then speak intomicrophone.

(5) Shutdown.1. Mode selector (VHF/AM/FM panel) -

OFF.2. POWER ON switch (Voice security

panel) - OFF.3-13. VOICE SECURITY SYSTEM TSEC/KY-58(PROVISIONS ONLY).

NOTEVoice security system TSEC/KY-58 may be installed in lieu ofvoice security TSEC/KY-28.Complete provisions are providedto install either the KY-28 or theKY-58 voice security system onthe pedestal extension (fig. 2-10).

The KY-58 voice security system provides secure(ciphered) two-way voice communications for the pilotand copilot in conjunction with the VHF/AM/FMcommand set, UHF command set and the backup VOWset. Voice security circuits are protected by the VHF,VHF/AM/FM, RADIO RELAY, and BU VOW circuitbreakers on the overhead circuit breaker panel fig. 2-26). For KY-58 operating instructions, refer to TM 11-5810-262-OP and TM 11-5810-262-20.

3-14. HF COMMAND SET (718 U-5).a. Description. The HF command set provides

long-range voice communications within the frequencyrange of 2.000 to 29.999 MHz and employs

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either standard amplitude modulation (AM), lowersideband (LSB), or upper sideband (USB) modulation.The distance range of the set is approximately 2,500miles and varies with atmospheric conditions. The unitis protected by a 3-ampere HF RCVR and the 25-ampere HF PWR circuit breaker on the overhead circuitbreaker panel (fig. 2-26). The control panel is locatedon the pedestal extension (fig. 2-10).

b. Controls and Indicators.(1) HF control panel fig. 3-7).

CONTROL/ FUNCTIONINDICATORFrequency Selects the desired operating fre-selector quency.Frequency Displays the selected frequency.

indicatorSQL control Adjusts squelch level.RF TEST Indicates operational status ofindicator set when RT TEST is selected

with the function selector.Function selector Turns set off and determines op-

erating mode.OFF Turns set off.USB Selects upper sideband modula-

tion.LSB Selects lower sideband modula-

tion.AM Selects amplitude modulation.CW Not used in this installation.SVU Not used in this installation.SVL Not used in this installation.RF TEST Check operational status of the

system.c. Normal Operation.

(1) Turn-on.1. Insure aircraft power is on.

NOTEIt is presumed the AVIONICSMASTER POWER switch is on andthat normally used avionic circuitbreakers remain depressed.

NOTE

Aircraft can be configured foreither HF or VOW on position 4 ofthe audio control panels.

2. AVIONICS MASTER POWER switch(overhead panel) - ON.

(2) Receive.1. HF audio switch (#4, audio panel) - ON.2. Function selector (HF panel, pedestal

extension) -Set to USB, LSB, or AM.3. Volume knob (audio switch #4, audio

panel) - Set to mid position.4. VOL knob (audio panel) - Adjust.5. SQL knob (HF panel) - Adjust to just

quiet noise when no signal is beingreceived.

6. Tuning knob (HF panel) - Set desiredfrequency.

(3) Transmit.1. Transmitter-interphone selector (audio

panel) -No. 4 position.2. Function selector (HF panel) - Set to

USB, LSB, or AM.3. MIC HEADSET/OXYGEN MASK switch

(instrument panel) -As desired.4. SQL knob (HF panel) - Adjust to just

quiet noise when no signal is beingreceived.

5. MIC switch - Depress to transmit.(4) Shutdown.

1. Mode selector knob (HF panel) - OFF.d. RF Test.

(1) Receive check.1. Set SQL control fully clockwise and tune

to WWV. Check receive operation.

NOTERadio station WWV operatescontinuously ton 2.500 Mhz,5.000Mhz, 10.000 Mhz, 15.000 Mhz,and 20.000 Mhz. Select thefrequency that provides the bestreceived signal strength.

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TM 55-1510-219-10

Figure 3-7.HF Control Panel (718 U-5)(2) RF TEST lamp check (system not keyed).

1. Set the function selector to RF TEST.Do not key the system. Observe the RFTEST lamp.a. If the RF TEST lamp does not light,

the lamp is defective or thereceiver-exciter is faulty.

b. If the RF TEST lamp blinks for morethan 1 minute, the receiver- exciteris faulty.

c. If the RF TEST lamp lightsimmediately or after an initial periodof blinking, the RF TEST lamp andfault circuits are operational.

d. Repeat RF TEST lamp check, asrequired, by selecting RF TEST withthe function selector.

(3) Tune check (system keyed).1. Set the function selector to RF TEST.

Change frequency selection to anyfrequency and key the systemmomentarily. Observe the RF TEST

lamp during the tune cycle. Normaltune time is 3 to 8 seconds.

NOTERF TEST lamp indications, duringtune, are only valid for 3 to 8seconds after initial key isapplied. To repeat the tunecheck, key the HF system with themicrophone switch.

a. If the RF TEST lamp stays oncontinuously, the receiver-exciter isfaulty.

b. If the RF TEST lamp blinks on andoff, the PA-coupler is faulty.

c. If the RF TEST lamp extinguishesafter a nominal tune time, andsidetone is heard when audio inputis supplied through the microphone,the system is tuned and operatingcorrectly for the selected frequency.

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e. HF Command Set - Emergency Operation.Not applicable.

3-15. EMERGENCY LOCATOR TRANSMITTER(ELT).

a. Description. An emergency locator transmitter(fig.3-8) provided to assist in locating an aircraft andcrew in the event an emergency landing is necessitated.The output frequency is 121.5 and 243 MHzsimultaneously. Range is approximately line of-sight.The transmitter unit has separate function controlswitches located on one end of the case. In the eventthe impact switch has been inadvertently actuated, thebeacon can be reset by firmly pressing the pushbuttonRESET switch on the front of the case. The RESETswitch and a 3-positiof toggle switch (placarded ARM,

OFF and ON) also on the transmitter case, may beactuated by inserting one finger through a small, round,spring-loaded door on the right side of the aft fuselage.The transmitter unit is accessible through a servicepanel located on the bottom of the aft fuselage.

b. Controls - ELT Transmitter Case.CONTROL FUNCTIONRESET switch When pressed, resets transmitterFunction switch Selects operating mode of set.ARM Arms set to be actuated by im-

pact switch (normal mode).OFF Turns set off.ON Activates transmitter for test

purposes.

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Figure 3-8. Emergency Locator Transmitter (ELT-10)

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TM 55-1510-219-10

Section III. NAVIGATION

3-16. DESCRIPTION.The navigation equipment group provides the pilot

and copilot with instrumentation required to establishand maintain an accurate flight course and position, andto make an approach on instruments under InstrumentMeteorological Conditions (IMC). The navigationconfiguration includes equipment for determiningattitude, position, destination range and bearing,heading reference and groundspeed.

3-17. RADIO MAGNETIC INDICATORS (RMI).a. Description. Two identical Radio Magnetic

Indicators (RMI) are installed, one for the pilot and onefor the copilot (fig. 3-9). Each unit serves as anavigational aid for the respective user and, by meansof individual source select switches will display aircraftmagnetic or directional gyro heading and VOR, TACAN,INS or ADF bearing information. The pilot's RMI isprotected by the 1-ampere #1 RMI circuit breaker on theoverhead circuit breaker panel (fig. 2-26) and the 1.5-ampere F13 fuse on the No. I junction box. Thecopilot's RMI is protected by the 1-ampere #2 RMI

circuit breaker on the overhead circuit breaker panel andthe 1.5-ampere F9 fuse on the No. 1 junction box.

b. Controls and Indicators.(1) Switches (fig. 2-28).

INDICATOR FUNCTIONPilot's Selects desired source of magnet-COMPASS No. ic heading information for dis-1/No. 2 switch play on pilot's HSI and copilot's

RMI.No. 1 Selects compass system No. 1

for display control.No. 2 Selects compass system No. 2

for display control.Copilot's Selects desired source of magnet-COMPASS No. ic heading information for dis-1/No. 2 switch play on co-pilot's HSI and pilot's


for display.No. 2 Selects compass system No. 2

for display.RMI select Selects which of two signals willswitch be displayed on respective RMI

single needle pointer, if single

Figure 3-9. Radio Magnetic Indicator (RMI) (332C-10)

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TM 55-1510-219-10

needle switch is in the VOR/TACAN position.VOR 1 Selects VOR 1 bearing signals

for display.TACAN Selects TACAN bearing signal

for display.(2) RMI (fig. 2-28).

CONTROL/ FUNCTIONINDICATORDouble needle Indicates bearing selected bypointer double needle switch.Compass card Indicates aircraft heading at top

of dial.Heading index Reference point for aircraft

heading.Warning flag Indicates loss of compass signal.Double needle Selects desired signal to be dis-switch played by double needle pointer.ADF position Selects ADF bearing informa-

tion.VOR position Selects VOR 2 bearing informa-

tion.Single needle Indicates bearing selected by sin-pointer gle needle switch.

Single needle Selects desired signal to be dis-switch played on single needle pointer.INS position Selects the bearing to waypoint

position.VOR/TACAN Selects VOR 1 or TACAN bear-position ing information.

3-18. HORIZONTAL SITUATION INDICATORS.a. Description. The pilot and copilot have separate

HSI instruments on respective instrument panelsections. Each HSI indicator combines displays toprovide a map-like presentation of the aircraft position.Each indicator displays aircraft heading, coursedeviation, and glideslope data. The pilot's HSI allowsthe desired course and autopilot input to be setmanually. Either HSI will display back localizer sensing,when the front course is selected and back course isflown. Course deviation is supplied to the HSI by theVOR 1 or VOR 2 systems, the TACAN (Tactical AirNavigation System), or the INS (Inertial NavigationSystem). Glideslope data is supplied by the VOR 1 orVOR 2 systems. The HSI displays warning flags whenthe VOR/LOC, INS, TACAN, heading or glideslopesignals are lost or become unreliable.

Figure 3-10. Pilot's Horizontal Situation Indicator (331A-8G)

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TM 55-1510-219-10

b. Controls and Indicators - Pilot's HSI.(1) Pilot's HSI (fig. 3-10).

CONTROL/ FUNCTIONINDICATORRotating heading The rotating heading dial, whichdial rotates with the aircraft through-

out 360 degrees, displays gyrostabilized magnetic compass in-formation. The azimuth ring isgraduated in five-degree incre-ments.

Heading marker Indicates heading selected byHDG knob.

Lubber line This pointed fixed index markaligns with the aircraft symbol toindicate aircraft magnetic head-ing on the compass card.

Course arrow Indicates VOR or TACAN radi-al course selected by COURSEknob.

HEADING flag Indicates that the heading infor-mation displayed is invalid andshould not be used.

COURSE Provides a digital readout of theindicator selected course.TO-FROM Two to-from pointers are situat-pointers ed 180 degrees apart. The one

which is visible points towardthe station along the selectedVOR/TACAN radial.

NAV flag The NAV flag indicates that theinformation derived from the se-lected navigational beacon is in-valid and should not be used.

COURSE The yellow course arrow is posi-control tioned on the heading dial by

the course knob to select a mag-netic bearing that coincides withthe desired VOR/TACAN radialor localizer course. Like theheading marker, the course ar-row rotates with the heading dialto provide a continuous readoutof course error to the computer.When one of the radio modes isselected, the vertical commandbar in the flight director will dis-play bank commands to inter-cept and maintain the selectedradio source.

Aircraft symbol A fixed aircraft symbol corre-sponds to the longitudinal axisof the aircraft and lubber line

markings. The symbol shows air-craft position and heading withrespect to a radio course and therotating dial.

Course deviation This bar represents the center-bar line of the selected VOR or lo-

calizer course. The aircraft sym-bol shows pictorially, actualaircraft position in relation tothis selected course.

HDG control The heading marker is posi-tioned on the rotating dial bythe heading knob and displayspreselected compass heading.The heading marker rotates withthe heading dial so the differ-ence between the marker and thefore lubber line index is theamount of heading error com-puter. In the heading mode, theflight director vertical commandbar will display the requiredbank commands to bring the air-craft onto and maintain the se-lected heading.

Course deviation In VOR operation, each dot rep-dots resents five degrees deviation

from centerline. In ILS opera-tion, each dot represents one de-gree deviation from centerline.In INS position, each dot repre-sents 3.75 NM linear deviation.

GS flag Indicates that the informationdisplayed by the glideslopepointer is invalid and should notbe used.

Glideslope Indicates deviation from correctpointer glideslope during ILS approach.Azimuth marks Azimuth marks are fixed at 45°

bearings throughout 360 degreesof compass card for quick refer-ence.

(2) Pilot's VOR Switches (fig. 2-28).CONTROL FUNCTIONPilot's course Selects desired source of data forindicator selector display on pilot's HSI and inputswitch to autopilot flight computer.VOR I Selects data from VOR I system.VOR 2 Selects data from VOR 2 system.TACAN Selects data from TACAN sys-

tem.INS Selects data from INS.

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TM 55-1510-219-10

Pilot's compass Selects desired source of magnet-No. 1/No. 2 ic heading data for display onswitch HSI compass card.No. 1 Selects data from No. 1 com-

pass.No. 2 Selects data from No. 2 com-

pass.c. Controls and Indicators - Copilot's HSI.

NOTEA PILOT SELECT annunciator,located on the copilot side of theinstrument panel, illuminateswhenever the setting of thecopilot's COURSE INDICATORswitch matches that of the pilot'sCOURSE INDICATOR switch. Inthis condition the copilot's HSI is"slaved" to the pilot's HSI, and thecopilot has no control overcourse select functions of theselected NAV receiver. The PILOTSELECT annunciator does notilluminate when INS is selected.

(1) Copilot's HSI (fig. 3-11).CONTROL/ FUNCTIONINDICATOR

Heading marker Indicates heading selected byHDG knob.

COMPASS flag Indicates that the heading infor-mation displayed is invalid andshould not be used.

Lubber line This pointed fixed index markaligns with the aircraft symbol toindicate aircraft magnetic head-ing on the compass card.

Course pointer Indicates VOR/TACAN radialcourse selected by COURSEknob.

Rotating heading The rotating heading dial, whichdial rotates with the aircraft through-

out 360 degrees, displays gyrostabilized magnetic compass in-formation. The azimuth ring isgraduated in five-degree incre-ments.

COURSE Provides a digital readout of theindicator selected course.VOR LOC flag Flag indicates that information

derived from the selected navi-gational beacon is invalid andshould not be used.

COURSE The yellow course arrow is posi-control tioned on the heading dial by

Figure 3-11. Copilot's Horizontal Situation Indicator (331A-6P)

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TM 55-1510-219-10

the course knob to select a magnetic bearing that coincides

with the desired VOR radial or localizer course. Like the heading marker, the course arrow rotates with the heading dial to provide a continuousreadout of course error to the computer. When one of the NAV modes is selected, the vertical command bar in the flight director will display bank commands to intercept and maintain the selected radio source.

Course deviation This bar represents the center-bar line of the selected VOR or lo-

calizer course. The aircraft sym-bol shows pictorially, actualaircraft position in relation tothis selected course.

Aircraft symbol A fixed aircraft symbol corre-sponds to the longitudinal axisof the aircraft and lubber linemarkings. The symbol shows air-craft position and heading withrespect to a radio course and therotating dial.

TO-FROM Two to-from pointers are situat-pointers ed 180 degrees apart. The one

which is visible points towardthe station along the selectedVOR radial.

HDG control The heading marker is posi-tioned on the rotating dial bythe heading knob and displayspreselected compass heading.The heading marker rotates withthe heading dial so the differ-ence between the marker and thefore lubber line index is theamount of heading error appliedto the flight director computer.In the heading mode, the flightdirector vertical command barwill display the required bankcommands to bring the aircraftonto and maintain the selectedheading.

Course deviation In VOR operation, each dot rep-dots. resents five degrees deviation

from centerline. In ILS opera-tion, each dot represents one de-gree deviation centerline. In INSoperation, each dot represents

3.75 NM deviation from center-line.

GS flag Indicates that the informationdisplayed by the glideslopepointer is invalid and should notbe used.

Glideslope Indicates deviation from correctpointer glideslope during ILS approach.Azimuth marks Azimuth marks are fixed at 45°

bearings throughout 360 degreesof compass card for quick refer-ence.

(2) Copilot's VOR Switches (fig. 2-28).Copilot's Course Selects desired source of data forIndicator display on copilot's HSI and in-Selector Switch put to autopilot.VOR 1 Selects data from VOR 1 system.VOR 2 Selects data from VOR 2 system.TACAN Selects data from TACAN sys-

tem.INS Selects data from INS.Copilot's Selects desired source of magnet-compass No. 1/ ic heading data for display onNo. 2 switch HSI compass card.No. 1 Selects data from No. 1 com-

pass.No. 2 Selects data from No. 2 com-

pass.

3-19. HORIZON REFERENCE INDICATOR.a. Description. The horizon reference indicator

(HRI) (fig. 3-12) is the pilot's basic attitude horizonindicator and the attitude direction instrument for theautomatic flight control system.

b. Controls and Indicators.CONTROL FUNCTIONCrossed needles Display computed commands to

autopilot.Lateral deviation Displays localizer deviation in-indicator formation from VOR receiver.Vertical Displays glideslope deviation in-deviation formation from VOR receiver.

indicatorBank angle Indicates aircraft bank angle.

pointerBank angle index Reference indicating zero-degree

bank.Bank angle scale Allows measurement of aircraft

bank angle from zero to 60 de-grees.

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Figure 3-12. Horizon Reference Indicator (329B-9A)

Horizon line Affixed to sphere, remains paral-lel to the earth's horizon at alltimes.

Miniature Indicates attitude of aircraftaircraft with respect to the earth's hori-zon.

Sphere Remains oriented with theearth's axis at all times.

GYRO flag Presence indicates loss of powerto, or low rotational speed of,vertical gyro.

CMPTR flag Presence indicates a malfunctionwithin the autopilot computer.

LOC flag Presence indicates that localizerinformation is not available ornot reliable.

NOTEWhen flying coupled to the INSsystem, the CMPTR flag will be inview anytime the steeringinformation is invalid or amalfunction exists in the autopilotcomputer.

GS flag Presence indicates glideslope in-formation is not being presentedon indicator.

TEST When pressed, display indicatespushbutton an additional 10° nose up, 20°right roll and the GYRO flag isvisible.

3-20. PILOT'S TURN AND SLIP INDICATOR.

a. Description. The pilot's turn and slip indicator(fig. 3-13) is used to provide automatic yaw dampinginformation to the autopilot in addition to performing thefunctions of a turn -and slip indicator. It is protected bythe 5-ampere PILOT TURN & SLIP circuit breakerlocated on the overhead circuit breaker panel (fig 2-26).

b. Controls and Indicators.INDICATOR FUNCTIONTurn rate Deflects to indicate rate of turn.indicatorTwo-minute turn Fixed markers indicate two-marks minute turn rate when covered

by turn rate indicator.GYRO warning Presence indicates loss of power

to instrument.

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Figure 3-13. Pilot's Turn and Slip Indicator (329T- 1)

Inclinometer Indicates lateral acceleration( side slip) of aircraft.

3-21. COPILOT'S GYRO HORIZON INDICATOR.a. Description. The copilot's gyro horizon indicator

(fig. 3-14) is a flight aid which indicates the aircraft'sattitude. The indicator is designed to operate through allattitudes. There are no front panel fuses or circuitbreakers provided for the copilot's gyro horizonindicator.

b. Controls and Indicators.INDICATOR FUNCTIONBank angle index Reference indicating zero-degree

bank.Bank angle Indicates aircraft bank angle.pointerBank angle scale Indicates aircraft bank angle

from zero to 90 degrees withmarks at 10, 20, 30, 45, 60, and90 degrees.

Horizon line Affixed to sphere, remains paral-lel to the earth's horizon at alltimes.

Miniature Indicates attitude of aircraftaircraft with respect to the earth's hori-

zon.G flag Presence announces loss of pow-

er.Sphere Indicates orientation with

earth's axis at all times.Inclinometer Assists the copilot in making co-

ordinated turns.

3-22. GYRO MAGNETIC COMPASS SYSTEMS.a. Description. Two identical compass systems

provide accurate directional information for the aircraftat all latitudes of the earth. As a heading reference, twomodes of operation are used: directional gyro (FREE)mode, or slaved (SLAVE) mode. in polar regions of theearth where magnetic heading references are notreliable, the system is operated in the FREE mode. Inthis mode, the system furnishes an inertial headingreference, with latitude corrections introduced manually.In areas where magnetic heading references arereliable, the system is operated in the SLAVE mode. Inthis mode, the directional gyro is slaved to the magneticflux detector, which supplies long-term magneticreference to correct the apparent drift of the gyro.Magnetic heading information from both systems is

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Figure 3-14. Copilot's Gyro Horizon Indicator (GH-14)

applied to various aircraft systems through pilot andcopilot COMPASS No. 1/No. 2 switches (fig. 2-28).There are no circuit breakers for the gyro magneticcompass systems. The circuits are protected by the 2-ampere F2 and F6 fuses on the No. 1 junction box.

b. Vertical Gyro. A vertical gyro provides line-ofsight stabilization to the weather radar and roll and pitchinformation to the autopilot. No controls are required orprovided for operation of the vertical gyro system. Thecircuit is protected by the 3ampere F22 fuse in the No.1 junction box.

c. Controls and Indicators.CONTROL/ FUNCTIONINDICATORPilot's Selects desired source for mag-COMPASS No. netic heading information to dis-1/No. 2 switch play on pilot's HSI and copilot's


for display.No. 2 Selects compass system No. 2

for display.Copilot's Selects desired source for mag-COMPASS No. netic heading information to dis-

1/No. 2 switch play on copilot's HSI and pilot'sRMI.

No. 1 Selects compass system No. 1for display.

No. 2 Selects compass system No. 2for display.

COMPASS Provides visual indication ofSLAVE system synchronization opera-annunciator tion.GYRO SLAVE/ Selects system mode of opera-FREE switch tion.SLAVE Selects slaved mode of opera-

tion.FREE Selects free mode of operation.INCREASE/ Provides manual fast synchroni-DECREASE zation of the system.

switchINCREASE Causes gyro heading output to

decrease (move in counter-clockwise direction).

DECREASE Causes gyro heading output toincrease (move in clockwise di-rection).

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d. Normal Operation.

(1) To align.

1. Compass GYRO SLAVE/FREEswitch SLAVE.

2. Compass INCREASE/ DECREASEswitch Hold switch momentarily inthe direction desired, and thenrelease. This will place system infast erect mode. The gyro will thenerect at approximately 30 degreesper minute. While in the fast erectmode, the HEADING flag (HSI) willbe in view. When the HEADINGflag retracts from view, the headingdisplayed will be the magneticheading.

(2) To determine magnetic heading.

1. Compass GYRO/SLAVE FREEswitch SLAVE.

2. RMI rotating heading dial (compasscard) Read heading.

(3) To determine directional gyro heading.

1. Compass GYRO/SLAVE FREEswitch-FREE.

2. Compass INCREASE/ DECREASEswitch Hold until the RMI compasscard aligns with the magneticheading, then release.

3. Read heading. The heading willagree with the appropriate HSI.

e. Shutdown. Both compass systems are shutdown when the inverter switch is turned off. (If eithersystem is ON, both compass sets will be energized.)

3-23. VOR/LOC NAVIGATION SYSTEM.

a. Description. The aircraft is equipped with twoVOR systems which are controlled by the NAV-I NAV-2control panel located in the pedestal (fig. 2-11).

Either VOR can direct input signals to the flightdirector indicator. Controls are shown in figure 315.

Each VOR system includes independent receiverunits for VOR/LOC and glideslope (GS). Each VORreceiver provides a VOR input to a respective RMIindicator and VOR and localizer data to the flightdirector. Each glideslope receiver sends GS flag andpointer deviation information to the flight director.

VOR/LOC indicators may be used for navigationduring manual control of the aircraft, or the autopilotmay be coupled to the VOR system, accepting VORinputs to the autopilot computer.

The pilot's unit (VOR 1) is a navigation radio systemwhich receives and interprets VHF OmnidirectionalRadio Range (VOR) signals, localizer (LOC) signals,glideslope signals, and marker beacon signals. Thesystem operates in a VOR/LOC frequency range of108.00 to 117.95 Mhz. Glideslope frequencies (329.15to 335.00 Mhz) are paired with LOC frequencies, andare automatically selected when LOC frequencies areselected. LOC frequencies are those frequenciesbetween 108.00 and 112.00 Mhz that end in odd tenths(108.1, 109.3, 109.5, etc.). The marker beacon receiveroperates at 75 Mhz and is not tuneable.

VOR 2 is similar to VOR 1 except VOR 2 cannotreceive or interpret marker beacon signals.

Each VOR system provides course deviation andglide path data, which can be switched either to thecopilot's HSI or to the autopilot flight computer andpilot's HSI, or both. The audio outputs of VOR 1 andVOR 2 systems are supplied to the NAV control, on theaudio control panels. VOR 1 bearing data is supplied tothe single needle pointer on both Radio MagneticIndicators. VOR 2 bearing data is supplied to thedouble-needle pointer on both Radio MagneticIndicators.

Both systems use a single VOR/LOC antenna,located on the vertical stabilizer, and a single glideslopeantenna, located in the radome. VOR 1 uses a markerbeacon antenna located on the underside of the forwardfuselage.

The VOR 1 system is protected by the 2-ampereVOR #1 and 35-ampere AVIONICS MASTER PWR #1circuit breaker on the overhead circuit breaker panel(fig. 2-26). The VOR 2 system is protected by the 2-ampere VOR #2 and 35-ampere AVIONICS MASTERPWR #2 circuit breakers on the overhead circuit breakerpanel.

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Figure 3-15. Dual NAV 1 - NAV 2 Control Panel (VIR-30AGM, VIR-30AG)

b. Controls and Indicators.

(1) NA V 1 Panel Section (VOR 1).CONTROL/ FUNCTIONINDICATORFrequency Displays selected frequency ofindicator VOR 1 receiver.Frequency Selects operating frequency ofcontrol VOR 1 receiver.VOL/OFF Activates VOR 1 receiver. Per-control mits monitoring VOR 1 audio

and adjust volume of signals re-ceived.

NAV TEST Activates self test of VOR 1 nav-switch igation system.

(2) NA V 2 Panel Section (VOR 2).CONTROL/ FUNCTIONINDICATORVOL/OFF Activates VOR 2 receiver. Per-control mits monitoring VOR 2 audio

and adjust volume of signals re-ceived.

Frequency Selects operating frequency ofcontrol VOR 2 receiver.

Frequency Displays selected frequency ofindicator VOR 2 receiver.NAV TEST Activates self test of VOR 2 nav-switch igation systems.

(3) Marker Beacon Indicators. (fig. 2-28).INDICATOR FUNCTION"A" indicators Illuminate when aircraft passes(white) over an inner marker beacon."O" indicators Illuminate when aircraft passes(blue) over an outer marker beacon."M" indicators Illuminate when aircraft passes(amber) over a middle marker beacon.

(4) Switches (fig. 2-28).CONTROL FUNCTIONPilot's COURSE Selects VOR receiver to controlINDICATOR pilot's HSI.switchVOR 1 VOR 1 controls pilot's HSI.VOR 2 VOR 2 controls pilot's HSI.Copilot's Selects VOR receiver to controlCOURSE copilot's HSI.INDICATORswitch

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VOR 1 VOR 1 controls copilot's HSI.VOR 2 VOR 2 controls copilot's HSI.NAV-A switch Applies VOR audio to respective

headsets.MKR BCN HI- Controls sensitivity of markerLO switch beacon receiver.

c. Operation.

(1) Turn-on.

1. Insure aircraft DC and AC power ison.

2. AVIONICS MASTER POWERswitch (overhead panel) ON.

3. Frequency controls (NAV panel) Setfor both receivers.

4. VOL knobs (NAV panel) Turnclockwise to activate sets, adjustvolume.

5. NAV A, audio switch ON. Confirmproper signal, then OFF.

6. Confirm proper indications RMI andHSI.

(2) Normal operation.

(a) Pilot, copilot COURSE INDICATORswitches (instrument panel): SelectVOR source.

(b) To determine course to station onpilot's HSI: Rotate course knob untilcourse deviation pointer is centeredand TO-FROM flag reads TO.

(c) To determine bearing from stationon pilot's HSI.: Rotate course knobuntil course deviation pointer iscentered and TO-FROM flag readsFROM.

(d) To determine course to station onRMI.: Select VOR, verify needlepoints course to station.

(3) Localizer (LOC) operation.

1. VOR frequency knob (NAV panel)Select frequency.

2. Pilot's, copilot's COURSEINDICATOR switches (instrumentpanel) Select VOR source.

3. Course deviation bar (HSI) Steeraircraft to center bar.

(4) Glideslope operation.

1. Frequency selectors (NAV panel)Set desired localizer frequency.

2. VOR switch (instrument panel)Select VOR source.

3. Glideslope pointer (HSI) Fly aircraftto center pointer.

(5) Self-test.

(a) NA V 1.

1. Frequency select knob (NAV panel)Select a VOR frequency.

2. NAV TEST switch (NAV panel)Press and hold.

3. RMI Observe that single needleindicates approximately 005°.

4. VOR/LOC flag Check that flag is outof view.

5. TO/FROM pointer Check thatpointer indicates TO.

6. HSI course deviation bar Check forcentered bar.

7. Marker beacon lights Check thatall three lamps are illuminated andflickering at approximately a 30 Hzrate.

8. VOR frequency knob (NAV panel)Select a LOC frequency.

9. VOR/LOC flag Check that flag is outof view.

10. HSI course deviation bar Check thatbar indicates a deflection ofapproximately one dot right ofcenter.

11. HSI glideslope pointer Check thatpointer indicates a deflection ofapproximately one dot below center.

12. Marker beacon lights Check thatall three lamps are illuminated andflickering at approximately a 30 Hzrate.

(b) NA V 2. All NAV 2 self-test procedures are the sameas those used for NAV 1, with theexception of the marker beacontest. There is no marker beaconreceiver in the NAV 2 system.

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(6) Shutdown. VOL knobs (NAV panel) OFF.

3-24. MARKER BEACON.

a. Description. A marker beacon receiver modulelocated inside the No. 1 VOR receiver, detects amarker beacon signal as the aircraft passes over amarker beacon transmitting antenna. The detectedsignal is selectively filtered to activate the appropriatemarker beacon lamp on the instrument panel. Inaddition, the detected signal is coupled to the aircraftaudio system to annunciate marker beacon passage.The marker beacon receiver operates on a fixedfrequency of 75 Mhz. Volume and sensitivity of thesystem are adjusted on the radio control panel (fig. 3-2)located on the pedestal extension (fig. 2-11).


CONTROL FUNCTIONMKR BCN HI- Controls sensitivity of markerLO switch beacon receiver.MKR BCN VOL Adjusts volume of marker bea-control con audio.

(1) Marker beacon indicators (fig. 2-28).INDICATOR FUNCTION"A" indicators Illuminate when aircraft passes(white) over an inner marker beacon."O" indicators Illuminate when aircraft passes(blue) over an outer marker beacon."M" indicators Illuminate when aircraft passes(amber) over an middle marker beacon.

c. Operation.

1. Marker beacon indicator lights(instrument panel) -Confirm beaconindication.

2. Marker beacon HI-LO switch (radiocontrol panel) -As required.

3. Select NAV A on audio controlpanel to monitor marker beaconaudio.

3-25. AUTOMATIC DIRECTION FINDER (DF-203).a. Description. The Automatic Direction Finder

(ADF) is a radio navigation system which provides avisible and audible indication of aircraft bearing relativeto a selected ground radio station. It may also be usedto home on a selected station, find

aircraft position, or monitor conventional low frequencyAM radio transmissions. The system is designed toprovide reliable reception of a 400-watt radio station at arange of 65 nautical miles throughout a 360-degree turnof the aircraft. It operates in a frequency range of 190to 1750 kilohertz.. Bearing indications are displayedvisually on the RMI's and aural signals are applied to theaudio control panels. Refer to Audio Control Panels (fig.3-1) for controls used to monitor ADF audio signals.

The ADF consists of a receiver, located on theforward side of the aft cabin bulkhead inside thepressure vessel; a control unit, located on the pedestalextension;, a non-directional sense antenna, installed inthe aircraft dorsal fin; a directional loop antenna, locatedon the underside of the fuselage; and a quadrangle errorcorrector, installed on the loop antenna (to compensatefor the deflection of arriving radio signals by the wingsand fuselage of the aircraft). The system is protected bythe 1-ampere ADF, the 5-ampere RADIO RELAY, andthe 35ampere AVIONICS BUS #2 circuit breakers onthe overhead circuit breaker panel (fig. 2-26).

NOTEKeying the HF radio set whileoperating the ADF set will cause amomentarily unreliable ADFsignal.


(1) ADF control panel (fig. 3-16).CONTROL/ FUNCTIONINDICATORLOOP control Operative only when the func-

tion switch is in the LOOP orADF position. Center positionremoves rotation signals fromthe loop antenna and the ADFpointer on the RMI's. First posi-tion L (left) or R (right) of centerapplies slow speed rotation sig-nals to loop antenna and ADFpointer on RMI's for 360-degreerotation left or right. Second po-sition L(left) or R (right) of cen-ter applies fast speed rotationsignals to loop antenna and ADFpointer on RMI's for 360-degreerotation left or right. Center po-sition holds antenna position.

BFO/OFF At BFO (on) setting, permitsfine tuning with Beat FrequencyOscillator (BFO). Also providesaudio tone when receiving un-modulated CW. OFF turns BFOoff.

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Figure 3-16. ADF Control Panel (DF-203)

Tuning meter Indicates relative strength of re-ceived signals.

TUNE control Selects operating frequency.Range switch Selects operating frequency

band.FREQUENCY Indicates selected frequency.indicatorMode selector Selects operating mode.OFF Turns set off.ADF Permits automatic direction

finding or homing operation.ANT Permits reception using sense

antenna.LOOP Permits audio-null homing and

manual direction finding opera-tions.

GAIN control Adjusts volume of received sig-nal.

(2) Switches - audio panel (fig. 3-1).

CONTROL FUNCTION

NAV B volume NAV B volume control appliescontrol ADF audio to aircraft audio sys-

tem.

(3) Switches - radio panel (fig. 3-2).CONTROL FUNCTIONFILTER V-OFF Selects whether voice filter willswitch be used with ADF audio.FILTER R-OFF Selects whether range filter willswitch be used with ADF audio.

c. Operation.

(1) To operate set as automatic directionfinder.

1. Mode selector - ADF.

2. BFO-OFF switch - BFO.

3. Range switch - Select.

4. TUNE control - Rotate for maximumreading on tuning meter and zeroBFO beat.

5. GAIN control - As required.

6. Double needle switches (RMI, fig. 3-9) - As required.

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7. Double needle on RMI Read courseto station.

(2) To operate set for sense antenna directionfinding: 1. Mode selector ANT.

2. Range switch Select.

3. TUNE control rotate for maximumreading on tuning meter.

4. GAIN control As required.

(3) To operate set for audio-null directionfinding:

1. Mode selector ANT.

2. BFO-OFF switch BFO.

3. Range switch select.

4. TUNE control Tune desired station.

5. GAIN control Adjust for minimumaudio output.

6. Double needle switches (RMI, fig. 3-9) As required.

7. BFO-OFF switch OFF.

8. Mode selector LOOP.

9. LOOP switch L or R. Turn left orright until a null is reached(minimum sound in headsets).

10. Double needle on RMI (fig. 3-9)Read course to station.

NOTEThe true null and direction to theradio station may be indicated byeither end of the single needle.This ambiguity must be solved todetermine proper direction to thestation.(4) Shutdown. Mode selector OFF.

3-26. TACAN SYSTEMS.

a. Description. Two Tactical Air Navigation(TACAN) systems are provided. One is dedicated to theINS and is used only for position updating; the other isused in conjunction with other avionics systems,including the flight director system and the autopilot.TACAN is a radio navigation system which providesaircraft distance and bearing information relative to aTACAN ground station. Both systems operate in the L-band frequency range of 962 to 1213 MHz.

Their range, though limited to line-of-sight, is designedto provide reliable reception of a TACAN ground stationat a distance of 170 nautical miles at an aircraft altitudeof 20,000 feet. The normal time required for thesystems to lock on to a selected ground station signal isthree seconds. The avionics TACAN system isprotected by the 2ampere TACAN circuit breakerlocated on the overhead circuit breaker panel (fig. 2-26).

b. Avionics TACAN System. The AN/ARN136(V)avionics TACAN system consists of the RT1321receiver-transmitter and the CP-1398 azimuth computer,both of which are located in the right nose avionicscompartment; the ID-2218 range indicator, located onthe instrument panel, the C-11265 control panel, locatedin the pedestal extension; and an antenna located onthe top surface of the aircraft fuselage. The avionicsTACAN system operates in conjunction with TACAN andVORTAC ground stations to provide distance, groundspeed, time-to-station, and bearing-to-station data. Itoperates in the L band frequency range on one of 252pre-selected frequencies, 126 X mode and 126 Y modechannels. Course deviation from TACAN stations isdisplayed on the HSI. Distance, time-to-station, andground speed are displayed on the TACAN digitaldisplay (fig. 3-17). The ground speed and time-to-station are accurate only if the aircraft is flying directlytoward the ground station at a sufficient distance thatthe slant range and ground range are nearly equal.

The avionics TACAN system may be operated bythe flight director system or connected to and used withthe autopilot system. When employed as the primarymeans of navigation, aircraft flight may be controlledmanually or by the autopilot. Indications of aircraftheading and bearing to ground stations are displayed onthe course deviation indicators (HSIs). Relative bearingto a station is displayed by the RMI bearing pointer.TACAN distance, ground speed, and time-to-station areall displayed on the TACAN indicator located on thecopilot's instrument panel (fig. 2-28).

The TACAN control panel (fig. 3-17) enablesselection of the TACAN frequency (channel) to be used,and provides for self-test of TACAN circuits. X or Ychannel is selected by the X/Y switch. Most TACAN andVORTAC stations are operated on the X mode. When Ymode stations are operational, air navigation charts willdesignate the Y mode stations. A toggle switch providessystem power ON/OFF control. Audio control isprovided by a rotary control placarded VOL.

c. INS TACAN System. The INS TACAN systemis of the same type as the avionics TACAN system.This set, inaccessible to pilot control, is coupled directly

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to INS circuits. It is dedicated only to updating the INS,is activated when the INS is operational, and iscontrolled only by the INS. The INS TACAN consists ofa range unit and a distance indicator, both located onthe INS equipment rack and both identical to theircounterparts in the avionics TACAN, and an antennalocated on the underside of the fuselage. No controlsare required or provided for the INS TACAN system.The system is protected by a circuit breaker on the INSjunction box.

d. Controls and Indicators.CONTROL/ FUNCTIONINDICATORTEST switch Activates system self-test.Channel Displays TACAN channel select-indicator ed.X-Y switch Selects X or Y mode for TA-

CAN channels.VOL control Adjusts TACAN volume.Channel Dual knob for manual selectionselectors of operating channel.Outer knob Selects tens and hundreds part

of channel number.Inner knob Selects units part of channel

number.

ON-OFF switch Activates or deactivates system.NM indicator Displays slant range distance in

nautical miles from aircraft toselected TACAN ground station.

KT indicator Displays ground speed in knots.MIN indicator Displays time to TACAN station

in minutes.e. Avionics TACAN Operation.

(1) Turn-On.

1. ON-OFF switch (TACAN panel) ON.

2. VOL knob As required.

3. Course indicator switches(instrument panel) -Select TACAN.

NOTEThe pilot and copilot should notselect TACAN simultaneously fortest.(2) Normal operation.

1. Pilot's single needle switchVOR/TACAN.

Figure 3-17. TACAN Control Panel and Range Indicator (ID-2218)

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2. RMI select switch (side panels)TACAN.

3. Pilot's Course indicator switchTACAN.

4. Mode switch (X/Y) As required.

5. TACAN control panel Select desiredchannel.

6. Wait 5 seconds for signalacquisition and lock-on.

7. Insure that audio stationidentification signal is correct for theground station selected.

8. Bearing pointer on RMI Readbearing heading to station.

9. COURSE knob (pilot's HSI) Setcourse desired.

10. Course deviation bar (pilot'sHSI)Read deviation from selectedcourse. Course arrow will showwind correction angle when thecourse deviation bar is centered andthe aircraft is tracking the selectedcourse.

11. TACAN indicator Read range (NM).

12. To determine course TO orcourse FROM a TACAN station,rotate course knob (pilot's HSI) untilcourse deviation bar is centered andthe TO/FROM flag reads TO orFROM.

13. To use TACAN during pilot-controlled flight, control aircraft bymanual controls, responding toinformation displayed on the flightdirector, RMI, TACAN, and otherinstruments.

14. To use TACAN with the autopilot,depress A/P ENGAGE and monitorautopilot performance on flightdirector, RMI, and TACANindicators. Verify adherence topreset heading and course, andconfirm the execution ofdisplayed steering commands.

NOTEThe TACAN ground speed readingwill be accurate only when theaircraft is on a course directly toor from the TACAN station. Whenheaded away from the station, theTACAN indicator minutes readingwill be in error.

(3) System test.

1. TEST pushbutton Press and hold.

2. Range indicator Check for anindication of 0.0 ± 0.1 nauticalmiles.

3. Pilot's COURSE selector switchSelect TACAN.

4. Pilot's RMI selector switch SelectTACAN.

5. RMI double needle Check for anindication of 180° ± 2°.

6. HSI course selector Turn to 180°and adjust slowly until the coursedeviation bar is centered. The barshould center between a selectedcourse of 178° to 182°.

7. HSI course selector Turn theselector + 10° from the settingachieved in step 6, and check thatcourse deviation bar is located overthe far left 10° dot.

8. HSI course selector Turn theselector -10° from setting achievedin step 6, and check that coursedeviation bar is located over the farright 10° dot.

9. TO-FROM indicator Check that TOis indicated.

10. TEST pushbutton Release.

(4) Shutdown. Turn ON/OFF switch OFF.

3-27. AUTOMATIC FLIGHT CONTROL SYSTEM.

a. Description. The AP-106 is an integratedautopilot/flight director system that provides thefollowing:

Heading mode

Navigation mode

Approach mode with automatic glideslopecapture and track

Altitude hold mode

Back course localizer mode

Go-around mode

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Synchronized control wheel steering

Indicated airspeed hold mode

All-angle adaptive capture for VOR, LOC, and LOC B/C

Attitude display

Heading display

Mode selection indicators

Elevator trim indicator

System integrity warning flags

Automatic yaw damping

Turn and slip indicator

The flight director and autopilot have a commoncomputer system. When the autopilot is engaged, theflight control system controls the aircraft and the pilotmonitors the flight path by observing the informationdisplayed on the pilot's horizon reference indicator (HRI)and the pilot's horizontal situation indicator (HSI) (flightdirector system indicators).

Autopilot/flight director commands are selected atthe autopilot mode selector panel (fig. 3-18) on thepilot's side of the instrument panel. Manual roll rate andpitch commands are inserted at the autopilot pitch-turnpanel (fig. 3-21). Autopilot operational status isindicated by the autopilot/flight director annunciatorpositioned above the pilot's horizon reference indicatoron the instrument panel (fig. 2-28). Two autopilotswitches are also provided on each control wheel (fig.2-16). One switch is placarded PITCH SYNC & CWS(pitch synchronize and control wheel steering), and theother is placarded DISC TRIM/AP YD (disconnecttrim/autopilot yaw damp).

Power for the system is provided through a10ampere AP PWR circuit breaker located on theoverhead circuit breaker panel (fig. 2-26).

b. Controls and Indicators. The following controlsand indicators are provided for operation of the system.

(1) Instrument Panel (Pilot's).

(a) Autopilot Mode select panel (fig. 3-18). The autopilot/flight director commands are selectedby the autopilot mode selector. Selection isaccomplished by pressing the face of the appropriatepush-on/push-off switch. The lateral modes are HDG,NAV, APPR and B/C. When not in a lateral mode, theflight director command bars are biased out of view.The vertical modes are ALT, IAS, and pitch (all holdmodes). If a vertical mode is not selected, the pitchhold mode is automatically operational.

Selection of a mode causes the legend of thatpushbutton switch to illuminate. The self-test switch onthe lower right of the autopilot control panel acts as alamp test when depressed. For operation at night,overall illumination of the autopilot mode selector andswitches is adjusted by the PILOT'S INSTRUMENT lightcontrol.

CONTROL/ FUNCTIONINDICATORHDG switch Engages heading mode. Com-

mands aircraft to acquire theheading indicated by headingmarker on pilot's HSI.

NAV switch Engages navigation mode. VOR-I/VOR-2 or TACAN selected,commands aircraft to interceptand track VOR radial selectedby course knob on pilot's HSI.INS selected, commands aircraftto track steering signals fromINS system. Intercept of approx-imately 45° and tracking will becomputed by the INS system.

NOTEAPPR cannot be selected with INS

selected.

APPR switch Engages approach mode. Com-mands aircraft to intercept andtrack ILS inbound course.

ALT switch Engages altitude hold mode.Commands aircraft to maintainpressure altitude.

IAS switch Engages indicated airspeed holdmode. Commands aircraft tomaintain airspeed.

B/C switch Engages backcourse mode. Com-mands aircraft to intercept backcourse ILS.

ENG-DIS switch Controls coupling of the auto-matic pilot.

ENG Engages autopilot and illumi-nates engaged indicator.

DIS Disengages autopilot and illumi-nates disengaged indicator.

TRIM UP Illuminates when autopilot isindicator driving trim servo in updirection.

TRIM DN Illuminates when autopilot isindicator driving trim servo in down di-

rection.

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Figure 3-18. Autopilot Mode Selector Panel (614E-42A)

Self-test switch Tests display and selector indi-cator circuits when depressed.

(b) Autopilot trim test switch (fig 3- 18).

AUTOPILOT Simulates trim system malfunc-TRIM TEST tions and illuminates AP TRIMswitch FAIL warning annunciator light.

(c) Aileron high torque mode test switch and annunciator(fig. 3-19).

CONTROL/ FUNCTIONINDICATORAIL HI Illumination is automatic fromTORQUE ground to 10,000 feet to showannunciator aileron servo is set to operate in

high torque mode. Light extin-guishes automatically above 10,000 feet to indicate aileron servohas terminated high torquemode operation.

TEST switch Normally off. Used only below10,000 feet (TEST position) toconfirm operability of aileronservo high torque mode, whenAIL HI TORQUE annunciatorlight extinguishes.

Figure 3-19. Aileron High Torque Test Switch andAnnunciator

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(d) Autopilot/flight director annunciatorpanel (fig. 3-20). The autopilot/flight directorincorporates its own annunciator panel located justabove the flight director display on the instrument panel.The modes and indications given on the annunciatorpanel are placarded on the face of the lenses andilluminate when the respective conditions are indicated.Dimming of the annunciator panel lights is provided by aswitch adjacent to the panel placarded DIM-BRT.

CONTROL/ FUNCTIONINDICATORNAV ARM Illuminates when computer isindicator armed to accept navigation sig-

nals.NAV CAP VOR-I/TACAN illuminatesindicator when selected radial is captured.

INS selected, illuminates whenINS is coupled to the flight di-rector.

GS ARM Illuminates when approachindicator mode is selected prior to glide-

slope capture. Extinguishes afterglideslope capture.

GS CAP Illuminates when glideslope isindicator captured.

GA indicator Illuminates when go-aroundmode is selected.

BACK LOC Illuminates when back courseindicator mode is selected.ALT indicator Illuminates when altitude hold

mode is selected.AP DISC Illuminates when autopilot isindicator disengaged.AP ENG Illuminates when autopilot is en-indicator gaged.IAS indicator Illuminates when airspeed hold

mode is selected.HDG indicator Illuminates when heading mode

is selected.LIN DEV Not ApplicableDIM BRT Adjusts intensity of illuminationcontrol of the flight director annunciator

(2) Pedestal - left power lever (fig. 2-11).CONTROL/ FUNCTIONINDICATORGO AROUND When pressed, autopilot discon-switch (outboard nects, GA annunciator light (fig.side left power 3-20) illuminates, and horizonlever) reference indicator commands

wings level, 7° nose up attitude.Autopilot may be re-engaged tofollow the command.

Figure 3-20. Autopilot/Flight Director Annunciator Panel

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(3) Pedestal extension.

(a) Autopilot pitch-turn panel (fig. 3-21).CONTROL FUNCTIONTurn control Supplies roll rate commands toknob autopilot. Spring loaded to cen-

ter detent.Pitch control Supplies pitch rate commands tothumbwheel autopilot. Spring loaded to cen-

ter detent.

(4) Control wheel switch (fig. 2-16).CONTROL/ FUNCTIONINDICATORDISC/TRIM/AP When pressed to first detent, au-YD pushbutton topilot system and yaw damp

are disconnected. When pressedto second detent, electric trim isdisconnected.

PITCH SYNC & This button, on each controlCWS pushbutton wheel, may be used instead of

the pitch-turn control to estab-lish the aircraft in a desired atti-

tude. Depressing the buttoncauses the autopilot servos todisengage from the control sur-

faces, enabling the pilot to man-ually fly the aircraft to the de-sired attitude until button isreleased.

c. Turn-on. Power is applied to the systemanytime the aircraft avionics bus is energized.

d. Autopilot Modes of Operation.

(1) Attitude. The autopilot is in the attitudemode when the ENG-DIS switch (AP mode select panel)is in the ENG position and no mode selector switches(HDG, NAV, etc.) have been selected (fig 3-18). Theautopilot will fly the aircraft and accept pitch and roll ratecommands from the autopilot pitch-turn panel (fig. 3-21).

(a) Guidance mode. When theautopilot is in the attitude mode and a mode selectorswitch (HDG, NAV, etc.) is pressed, the autopilotaccepts steering commands from the computer.Depending on which selector switch (AP mode selectpanel) is pressed, autopilot operation can be describedby the following subguidance modes.

(b) Heading mode. When HDG modeis selected (AP mode select panel), with au topilotengaged, the autopilot will fly the aircraft to, and thenmaintain, the heading under the heading marker on thepilot's HSI.

Figure 3-21. Autopilot Pitch-Turn Control Panel

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(c) Navigation mode. When NAV modeis selected (AP mode select panel), the system initiallyswitches to the NAV ARM heading hold submode, asshown by illumination of NAV ARM and HDG indicators(AP/flight director annunciator panel). The autopilot willthen command the aircraft to follow the heading underthe heading marker on the pilot's HSI (with the headingmarker set to produce the desired VOR or localizerintercept angle). The flight computer will compute acapture point based on deviation from the desired radiobeam, the rate at which the aircraft is approaching thisbeam, and the course intercept angle. When beamcapture occurs, the HDG and NAV ARM indicator lamps(AP/flight director annunciator panel) will extinguish andthe NAV CAP lamp will illuminate. The autopilot willthen track the selected radio course with automaticcrosswind correction.

(d) Back-course mode. When BACKLOC mode is selected on the autopilot mode selectorpanel, localizer capture is the same as in a frontcourseapproach in NAV or APPR mode. Glideslope isinhibited during a back-course approach. The HSI mustbe set to the front-course heading so that lateraldeviation will be directional.

(e) Approach mode. When APPRmode is selected (AP mode select panel), localizercapture is the same as in the NAV mode but glideslopearm and capture functions are also provided. When theAPPR mode is selected the NAV ARM annunciator lampwill illuminate, indicating that the system is armed forlocalizer capture. As the aircraft approaches thelocalizer beam, the NAV CAP annunciator lamp willilluminate. Once the localizer is being tracked, the GSARM annunciator lamp will illuminate. Glideslopecapture is dependent on localizer capture and mustoccur after localizer capture. The localizer is alwayscaptured from a selected heading, but the glideslopemay be captured with the autopilot operating in anyvertical mode (pitch hold, altitude hold, or indicatedairspeed hold), and from above (not recommended) orbelow the glideslope. At the point of glideslopeintercept, the GS CAP annunciator lamp will illuminateand all preselected vertical modes will be cleared.

(f) Go-around mode. Pressing the GOAROUND button on the outboard side of the left powerlever selects the go-around mode. Go-around modemay be selected from any lateral mode (HDG, NAV,APPR, or BACK LOC). When go-around mode isselected: (1)the autopilot is disengaged, (2) the GAannunciator lamp will illuminate, and (3) a commandpresentation for wings level and 3-38 7° nose up pitchattitude will appear on the pilot's horizon referenceindicator.

NOTEThe heading marker may bepreset to the go-around headingafter the localizer is captured.After go-around airspeed andpower settings are established,select the HDG mode to clear thego-around mode. Pitch attitudewill remain at that used for go-around until changed with thePITCH SYNC & CWS button or theselection of a vertical mode.

(g) Pitch hold mode. The pitch holdmode is selected by (1) selecting one of the verticalmode selector switches, or (2) actuating the pitchsynchronize and control wheel steering switch (PITCHSYNC & CWS), located on each control wheel.

(h) Control wheel steering mode.Pressing one of the PITCH SYNC & CWS switcheslocated on each control wheel disconnects the autopilotservos from the control surfaces, allows the pilot to flythe aircraft to a new pitch attitude, and synchronizes thevertical command bar (pilot's horizon ref ind) to aircraftattitude. The ALT or IAS mode will disengage (ifselected) when the PITCH SYNC & CWS button isdepressed. When the autopilot is coupled to the HDG,NAV, APPR, or BACK LOC modes, releasing the PITCHSYNC & CWS switch will cause the autopilot to coupleto the previously selected mode.

(i) Altitude hold mode. Pressing theALT switch (AP mode select panel) when desiredaltitude has been reached (with autopilot engaged) will(1) cause the autopilot to fly the aircraft to maintain thebarometric altitude at which the aircraft was flying whenALT switch was pressed, (2) illuminate the ALTannunciator lamp (AP/flight director annunciator panel),and (3) display the altitude hold commands on thevertical command bar (pilot's horizon referenceindicator)

(j) Indicated airspeed hold mode.Pressing the IAS switch (AP mode select panel) whendesired airspeed has been reached (with autopilotengaged) will (1) cause the autopilot to fly the aircraft tomaintain the indicated airspeed at which the aircraft wasflying when IAS switch was pressed, (2) illuminate theIAS annunciator lamp (AP/flight director annunciatorpanel), and (3) display IAS hold commands on thevertical command bar of the pilot's horizon referenceindicator.

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e. Autopilot Operation.

(1) Engaging autopilot. AP mode selectpanel ENG/DIS switch to ENG position.

NOTEWhen the autopilot is engaged,the yaw damper is alsoautomatically engaged asindicated by the lighted YAWDAMP button (AP pitch/turncontrol panel).

The autopilot and flight director are coupled whenboth units are engaged. When coupled, the autopilotaccepts guidance commands from the flight director.When the flight director is not engaged, the autopilotaccepts pitch and roll commands from the pitch/turncontrol knobs as selected by the pilot.

The autopilot may be engaged in any reasonableattitude and in either the coupled or uncoupled mode.The autopilot will smoothly acquire the commandattitude. When uncoupled, the autopilot will maintainthe bank and pitch attitude at the time of engagement.

(2) Disengaging autopilot.

(a) The autopilot may be disengaged byany of the following:

1. Actuating compassINCREASE/DECREASE switch.

2. Pressing TEST button on AP modeselect panel.

3. Pressing GO AROUND switchon left power lever.

4. Pressing control wheel DISC switchto AP YD position.

(b) The following functions will causethe autopilot to automatically disengage:

1. Vertical gyro failure.

2. Directional gyro failure.

3. Autopilot power or circuit failure.

4. Torque limiter failure.

(3) Maneuvering.

1. To change flight functions, press thedesired mode button on the APmode selector panel. The buttonwill illuminate along its edges andautopilot annunciator lights on theinstrument panel will illuminate,indicating the respective modes inoperation.

2. In any function except "afterglideslope capture", use theautopilot pitch control for climbingor descending. Movement of thepitch control establishes a pitch ratethat is proportional to knobdisplacement. If any verticalmode button has been selected, itwill automatically release when thepitch control knob is rotated.

3. When HDG mode is selected, theautopilot will command the aircraftto execute a turn then maintain theheading set by the heading marker.

4. Use the autopilot turn control tocommand a roll rate when theautopilot is engaged. At the timecontrol is returned to detent, theautopilot maintains the bank angle(up to approximately 30 degrees).Rotating the turn control when theautopilot is engaged and a lateralmode is selected will cause theselected lateral modes to release.

(4) Control wheel synchronization. ThePITCH SYNC & CWS button on the pilot's control wheelcan be used instead of the pitch/turn control to establishthe aircraft in a desired attitude. Depressing this buttoncauses autopilot elevator and aileron servos todisengage from the control surfaces. The pilot then fliesthe aircraft manually to a desired attitude, releases thePITCH SYNC & CWS button to re-engage the servos,and the autopilot holds the established attitude.

The ALT and IAS mode will immediately disengage(if selected) when the PITCH SYNC & CWS button ispressed. Upon release of the PITCH SYNC & CWSbutton, the autopilot will couple to the previouslyselected lateral mode.

NOT The APPR mode will not

disengage when the PITCH SYNC& CWS button is depressed.When the button is released, theaircraft will return to the localizercourse and glideslope.

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f. Takeoff and Climb Out.

(1) Before takeoff

1. Heading marker (pilot's HSI) Set torunway heading.

2. HDG switch (AP mode selectpanel) Press. Do not engageautopilot.

(2) Takeoff Pressing the PITCH SYNC &CWS switch on control wheel will provide pitch sync,and the cross-pointers on the pilot's horizon referenceindicator will command flight to the pitch attitude thatexisted when the PITCH SYNC & CWS switch waspressed.

(3) Climb out.

1. Establish climb profile.

2. ENG/DIS switch (AP mode selectpanel) Set to ENG (above 200 feetAGL).

3. IAS switch (AP mode select panel)Press (if desired).

4. HDG knob (pilot's course indicatorselector) Move heading marker asrequired for heading changes.

(4) Cruise altitude.

1. Vertical speed Reduce to approx.500 feet per minute (just beforereaching cruise altitude).

2. ALT button (AP mode select panel)Press (when reach cruise altitude).

g. VOR Operation.

(1) To establish aircraft on a desired VORradial, perform the following

1. VOR receiver Tune appropriatefrequency.

2. COURSE knob (pilot's HSI) Setdesired course TO or FROM stationshown in COURSE window.

3. HDG knob (pilot's HSI) Set desiredbeam intercept angle under headingmarker. (The intercept angle withrespect to the radio beam may beany angle of 90° or less).

4. NAV switch (AP mode select panel)Press. (Observe illumination ofNAV ARM annunciator).

5. NAV CAP annunciator (AP/flightdirector annunciator panel) Monitor.(At point of capture, NAV CAPilluminates).

NOTEExcept as described below, donot select a different VORfrequency, TACAN channel, orcourse once a course andintercept have been programmedor capture achieved. To select adifferent course or VOR/TACANfrequency, return to the HDGmode, select the course orfrequency, return to the NAVmode, then reset the desiredbeam.(2) To change course over a VOR station. To

change course over a VOR station while operating inNAV mode, if course change is less than 30°: COURSEknob (pilot's HSI) Set desired heading in COURSEwindow.

(3) To change course over a VOR station. Tochange course over a VOR station while operating inNAV mode, if course change is greater than 30°:

1. HDG knob (AP mode select panel)Set desired intercept heading underheading marker.

2. HDG switch (AP mode select panel)Press. (Observe HDG illuminates(AP/flight director annunciatorpanel).

3. COURSE knob (pilot's HSI) Set newcourse in COURSE window.

4. NAV switch (AP mode select panel)Press (Observe NAV ARMilluminates (AP/flight directorannunciator panel).

5. NAV CAP (AP/flight directorannunciator panel) -Monitor.(Illumination means capture of newradial).

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h. Automatic Approach Front Course.NOTE

The localizer and glideslope are capturedautomatically on the ILS front courseapproach. The localizer must be capturedbefore glideslope capture can occur. Thelocalizer is always captured from a selectedheading, but the glideslope may be capturedfrom any of the vertical modes and fromabove (not recommended) or below theglideslope.

1. VOR receiver Tune appropriatefrequency.

2. COURSE knob (pilot's HSI) Setinbound runway heading inCOURSE window.

3. HDG knob (pilot's HSI) Set headingmarker to desired intercept angle.

4. HDG selector switch (AP modeselect panel) Press. (Observe HDGilluminates on AP/flight directorannunciator panel).

5. Vertical mode (AP mode selectpanel) Select IAS or ALT.

6. APPR switch (AP mode selectpanel) Press. (Observe NAV ARMilluminates on AP/flight directorannunciator panel).

7. NAV CAP (AP/flight directorannunciator panel) -Observeillumination (when capture oflocalizer course is achieved).

8. GS ARM (AP/flight directorannunciator panel) -Observeillumination (when autopilot isarmed for glideslope capture).

9. GS CAP (AP/flight directorannunciator panel) -Observeillumination (which confirms that allvertical modes are cleared, and alsoconfirms that autopilot is trackingglideslope).

i. Go-Around. If visual runway contact is notmade at decision height, go-around may be activated bypressing the GA button on the left power lever, and maybe initiated from any lateral mode (HDG, NAV, APPR,B/C) with the following results:

1. Illuminates GA light on autopilotannunciator panel.

2. Disengages autopilot.

3. Pilot's HRI shows commandpresentation for wings level and 7°nose up climb attitude.

NOTEThe heading marker may be preset to go-around heading after the localizer iscaptured. After go-around airspeed andpower settings are established, selection ofthe HDG mode will clear the go-aroundmode. Pitch attitude will remain at that usedfor go-around until changed with the PITCHSYNC & CWS button or by selection of avertical mode.j. Back Course Approach. As in front course

approach, the localizer is captured automatically. Theaircraft should be maneuvered into the approach areaby setting the heading marker and functioning in theHDG mode.

1. VOR receiver (NAV panel) Tunelocalizer frequency.

2. COURSE knob (pilot's HSI) Setfront course inbound runwayheading in COURSE window.

3. HDG knob (pilot's HSI) Set headingmarker to desired intercept angle.

4. HDG switch (AP mode select panel)Press. (Observe HDG lampilluminates on AP/flight directorannunciator panel).

5. B/C switch (AP mode select panel)Press. (Observe BACK LOC andNAV ARM lamps illuminate(AP/flight director annunciatorpanel) indicating system is armedfor back localizer capture.

6. NAV CAP lamp will illuminate whensystem has captured back localizercourse.

7. PITCH control (AP/pitch-turn panel)Use to establish and maintaindesired rate of descent.

NOTEThe HDG mode should be used within onemile of the runway due to the large radiodeviations encountered when flying over thelocalizer transmitter.

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k. Yaw Damper Operation.

1. The rudder channel of the autopilotmay be selected separately for yawdamping by depressing the YAWDAMP switch on the pedestal. Theswitch face will illuminate when theyaw damper is engaged.

2. To disengage the yaw damper,press the disconnect button on thepilot's or copilot's control wheel tothe first detent or press the YAWDAMP switch on the pedestal.

3. Refer to Emergency Procedures forother means of disconnecting theyaw damper.

I. Disconnecting Autopilot. The autopilot may bedisconnected by any of the following actions:

1. Pressing the DISC TRIM/AP YDswitch to first detent. (Location:outboard horn either control wheel.)

2. Placing the ENG-DIS switch to DISposition. (Location: AP mode selectpanel.) NOTE After assumingmanual control, fly the aircraft usingthe same heading, course, andattitude displays to monitor autopilotoperation prior to assuming manualcontrol.

m. Emergency Procedures. The autopilot can bedisengaged by any of the following methods:

1. Press the AP/YD disconnect switch(either control wheel).

2. Move the engage lever to DISposition (either control wheel).

3. Engage the go-around mode (yawdamper will remain on).

4. AP PWR and AFCS DIRECT circuitbreakers (overhead panel) Pull.

5. AVIONICS MASTER POWERswitch (overhead panel) -OFF.

6. Aircraft MASTER switch (overheadpanel) OFF.

n. Automatic Disengagement. The followingconditions will cause the autopilot to disengageautomatically.

1. Any interruption or failure of power.

2. Vertical gyro failure indication.

3. Flight control system power orcircuit failure.

4. Autopilot trim failure.

3-28. INERTIAL NAVIGATION SYSTEM.

a. Description. The Inertial Navigation System(INS) is a self-contained navigation and attitudereference system. It is aided by (but not dependentupon) data obtained from the ground-based data linksystem, its own TACAN system, the aircraft encodingaltimeter, the true airspeed computer, and the gyromagnetic compass system. The position and attitudeinformation computed by the INS is supplied to theautomatic flight control system, weather radar system,horizontal situation indicator, radio magnetic indicators,and to mission equipment. The INS is independent ofaircraft maneuvers, weather conditions and terrain, andin conjunction with other aircraft equipment permitsoperations on instruments alone under InstrumentMeteorological Conditions (IMC). The INS provides avisual display of present position data in UniversalTransverse Mercator (UTM) coordinates or conventionalgeographic (latitude-longitude) coordinates during allphases of flight. When approaching the point selectedfor a leg switch, an ALERT lamp will light informing thepilot of an eminent automatic leg switch or the need tomanually insert course change data. The INS may bemanually updated for precise aircraft present positionaccuracy by flying over a reference point of knowncoordinates. The INS may be updated automatically bythe TACAN system or the Data Link system. Altitudeinformation is automatically inserted into the INScomputer by an encoding altimeter whenever the INS isoperational.

The Mode Selector Unit (MSU) (fig. 3-22) controlssystem activation and selects operating modes.

The Control Display Unit (CDU) (fig. 3-23) providescontrols and indicators for entering data into the INS anddisplaying navigation and system status information.

The INS system is protected by the 10-amperePRIME POWER and the 5-ampere HEATER POWERcircuit breakers on the mission AC/DC power cabinet, bythe 5-ampere INS CONTROL circuit breaker on theoverhead circuit breaker panel and by the 20-amperecircuit breaker on the front of the battery unit.

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Figure 3-22. Mode Selector Unit (C-IV-E)


(1) Mode selector unit (fig. 3-22).CONTROL FUNCTIONSMode select Controls INS activation and se-knob lects operating modes.OFF Deactivates INS.STBY (ground To STBY from OFF mode.use only) Starts fast warmup of system to

operating conditions; activatescomputer so information may beinserted; all INS-controlledwarning flags will indicate warn-ing.

To STBY from any other mode:

INS operates as if in attitude reference mode; allINS warning flags and lamps, except ATTITUDE andPLATFORM HEADING, will indicate warning.

ALIGN (ground To ALIGN from STBY mode.use only, parked) Starts automatic course align-

ment mode (if fast-warmup heat-ers are off); fine alignment willnot start until present position isinserted into CDU.

To ALIGN from OFF mode:

Leveling starts after fast-warmup heaters are off.

To ALIGN from NAV mode:

INS is not downmoded but will allow automaticshutdown if overtemperature is detected.

NAV Activates normal navigationmode after automatic alignmentis completed; must be selectedbefore moving aircraft.

To NAV from STBY mode:

Causes INS to automatically sequence throughSTBY and ALIGN to NAV mode, if present position isinserted and aircraft is parked.

Used to shorten time in STBY and to bypass batterytest, if stored heading is valid.

ATT Activates attitude referencemode. Used to provide only INSattitude signals. Shuts downcomputer and CDU leaving onlyBAT and WARN lamps opera-tive. Once selected, INS align-ment is lost.

BAT lamp Lights to indicate INS shutdowndue to low battery unit voltage.

READY NAV Lights to indicate INS high accu-lamp racy alignment is attained. If at-

tained during ALIGN mode,

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lamp remains illuminated untilNAV mode is selected.Lamp lights momentarily duringalignment, if alignment accom-plished while in NAV mode.

(2) INS control display unit 6fig. 3-23).CONTROL FUNCTIONSINDICATORHOLD key Used with other CDU controls

to stop present position displayfrom changing, in order to up-date position and to display re-corded malfunction codes.Lights when pressed first time;goes out when pressed secondtime or when inserted data is ac-cepted by computer.When pressed second time, al-lows displays to resume showingchanging current present posi-tion.

ROLL LIM key Allows selection of Limited Rollsteering mode. Press to selectmode, key lights. Roll steeringoutput is limited to 10 degrees.

Press second time to exit mode,key light extinguishes. Roll steer-ing output returns to normallimit of 25 degrees.

Data display, left Composed of lamps which illu-and right minate to display numbers, deci-

mal points, degree symbols, leftand right directions, and latitudeor longitude directions.

INSERT/ Allows insertion of loaded dataADVANCE/ HI into computer. Enters displayedPREC key data into INS. When pressed be-

fore pressing any numerical key,alternates display of normal andhigh precision data.

ALERT lamp Illuminates amber to alert pilot1.3 minutes before impendingautomatic course leg change. Ex-tinguishes when switched to newleg, if AUTO/MAN switch is setto AUTO.Flashes on and off when passingwaypoint, if AUTO/MAN switchis set to MAN. Light will extin-guish if AUTO is selected or if acourse change is inserted.

WARN lamp Lights red to alert pilot INS self-test circuits have detected a sys-tem fault. Illumination may be

Figure 3-23. INS Control Display Unit (C-IV-E)

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caused by continuous or inter-mittent condition. Intermittentconditions light lamp until resetby TEST switch. If continuouscondition does not degrade atti-tude operation, lamp goes outwhen mode selector is set toATT.

Keyboard Consists of 10 keys for enteringload data into data and FROM-TO displays.

"N", "S", "E", and "W" (on keys 2,8,6 and 4) indicatedirection of latitude and longitude. TAC and DISP (onkeys "7" and "9") enable loading and display of TACANstation data. MV and DISP (on keys "3" and "9") areassociated with loading and display of MagneticVariation and Magnetic Heading. GRID and DISP (onkeys "5" and "9") are associated with loading and displayof UTM coefficients.

CLEAR key When pressed, illuminates anderases data loaded into data dis-plays or FROM-TO display.Used to cancel erroneous data.After clearing, data loading canbe resumed.

WYPT CHG key When pressed, enables numbersin FROM-TO display to bechanged. If INSERT/ADVANCEkey is pressed, computer will usenavigation leg defined by newnumber in all navigation compu-tations. If INSERT/ADVANCEkey is not pressed, computer willcontinue using original numbersin all navigation computations;but distance/time information,based on new leg, may be calledup and read in data displays (incase of waypoints). When not inTACAN mix mode, TACAN sta-tion number is inserted to dis-play DIS/TIME information.

AUTO-MAN This is a dual purpose control.TEST switch When the knob is pressed in-

ward, the TEST switch functionis engaged. When the knob is ro-tated to either the AUTO oras a selector between thosemodes.

(AUTO) Selects automatic leg switchingmode. Computer switches fromone leg to the next wheneverwaypoint in "TO" side of theFROM-TO display is reached.

(MAN) Selects manual leg switching

mode. Pilot must makewaypoint changes manually.

(TEST) When pressed, performs test ofINS lamps and displays, remotelamps and indicators controlledby INS, and computer input/output operations.

Used with other controls to activate display ofnumerical codes denoting specific malfunctions andresets malfunction warning circuits.

During alignment, activates the HSI test. Continuedpressing of switch provides constant INS outputs todrive cockpit displays in a predetermined fashion.

NOTEThe INS can provide test signals to theHorizontal Situation Indicator (HSI)andconnected displays. Pressing TEST switchduring Standby, Align, or NAV modes willcause all digits on connected digitaldisplays to indicate "8's" and illuminates theHSI and ALERT lamps. Additional HSI testsignals are provided when INS is in Alignand the data selector is at any position otherthan DSRTK/STS. Under those conditions,pressing TEST switch causes the HSI toindicate heading, drift angle, and track angleerror all at "0°" or "30°". At the same time,cross track deviation is indicated at "3.75"nautical miles (one dot) right or left and INS-controlled HSI flags are retracted from view.

Output test signal are also supplied to the autopilotwhen INS steering is selected. Rotating AUTO/MANswitch to AUTO and pressing TEST during Alignfurnishes a 15° left bank steering command. A 15° rightbank steering command is furnished when theAUTO/MAN switch is set to MAN.

FROM-TO Display numbers definingdisplay waypoints of navigation leg be-

ing flown or in the case of aflashing display, displays TA-CAN station being used.

Waypoint numbers automatically change eachtime a waypoint is reached. Unless flight planchanges during flight, the automatic leg switching se-quence will always be 1 2, 2 3, 3 4....8 9, 9 1, 1 2,etc.Data selector Selects data to be displayed in

data displays or entered intoINS. The rotary selector has 10positions. Five positions (L/L

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POS, L/L WYPT, UTM POS,UTM WYPT and DSRTK/STS)also allow data to be loaded intodata display then inserted intocomputer memory.

TK/GS Displays aircraft track angle inleft display and ground speed inright display.

HDG/DA Displays aircraft true heading inleft display and drift angle inright display.

XTK/TK Displays cross track distance inleft display and track angle errorin right display.

L/L POS Displays or enters present air-craft position latitude in leftdata display and longitude inright data display. Both displaysindicate degrees and minutes tonearest tenth of a minute.

This position also enables the insertion of presentposition coordinates during alignment and presentposition updates.

L/L WY PT Displays or enters waypoint andTACAN station data, if used inconjunction with the waypoint/TACAN selector.

This position will also cause display of inertialpresent position data when the HOLD key is illuminated.

DIS/TIME Displays distance from aircraftto TACAN station or anywaypoint, or between any twowaypoints in left display. Dis-plays time to TACAN station orany waypoint, or between anytwo waypoints, in right display.

WIND Displays wind direction in leftdisplay and wind speed in rightdisplay, when true airspeed isgreater than the Air Data Systemlower limit (115-400 KIAS).

DSRTK/STS Displays desired track angle tonearest degree in the left datadisplay, and INS system statusin right data display.

UTM POS Displays or enters aircraft posi-tion in Universal TransverseMercator (UTM) coordinates,with northing data in kilometersin left display and easting datain kilometers in right display.

The extra precision displayshows meters.

UTM WYPT Displays or enters waypoint andTACAN station data in UTMcoordinates. Also enables load-ing and display of spheroid coef-ficients if GRID and DISP keysare pressed simultaneously.

Dim knob Controls intensity of CDU keylamps and displays.

Waypoint/DME Thumbwheel switch, used to se-selector lect waypoints for which data is

to be inserted or displayed.Waypoint station "O" is for dis-play only and cannot be loadedwith usable data.

c. INS - Normal Operating Procedures.

NOTEThe following data will be required prior tooperating the INS: Local magnetic variation,latitude and longitude or UniversalTransverse Mercator (UTM) coordinates ofaircraft during INS alignment. Thisinformation is necessary to program the INScomputer during alignment procedure.

NOTEWhen inserting data into INS computer,always start at left and work to right. Thefirst digit inserted will appear in rightposition of applicable display. It will step toleft as each subsequent digit is entered. Thedegree sign, decimal point, and colon (ifapplicable) will appear automatically.

(1) Preflight procedures.CAUTION

Insure that cooling air is available tonavigation unit before turning the INS on.

NOTEThe INS requires an aircraft compass systemto provide accurate navigational data andmission data. The compass system selectedwith the copilot's compass select switch willprovide information to the INS platform.

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NOTEAircraft must be connected to aground power unit if INSalignment is performed prior toengine starting. In this event, theengines must not be started untilafter the INS is placed in the NAVmode.

1. Applicable circuit breakers Checkdepressed.

2. INS Mode selector (MSU) ALIGN.Confirm following:

a. (CDU) FROM-TO display indicates"1 2".

b. (CDU) INSERT/ADVANCEpushbutton light illuminates.

c. (CDU) BAT lamp illuminates forapprox. 12 seconds, thenextinguishes when in alignmentstate 8.

NOTEAvoid passenger or cargo loadingor any activity which may causeaircraft to change position orattitude during alignment. Ifaircraft is moved duringalignment, it will be necessary torestart alignment by setting themode selector to STBY, then backto ALIGN and reinserting presentposition data.

3. (CDU) DIM knob Adjust for optimumbrightness of CDU displays.

4. (CDU) AUTO/MAN/TEST switchAUTO.

5. (CDU) Data selector L/L POS orUTM POS, as desired.

(Observe coordinates of last present position prior toINS shutdown appear in data displays.) NOTE Thestored heading procedure should have been performedwithin the last week.

6. Perform recommended storedheading alignment preflightprocedure:

a. Appropriate circuit breakersDepress (including those formagnetic compass system.)

b. (MSU) Mode selector NAV.

(Observe (CDU) FROM-TO display indicates "1 2",and INSERT/ADVANCE pushbutton illuminates.)

NOTEAircraft must not be towed ortaxied during INS alignment.Movement of this type duringalignment causes large navigationerrors. If aircraft is moved duringalignment, restart alignment bysetting mode selector to STBY,then back to ALIGN andreinserting present position.

Passenger or cargo loading in the aircraft couldcause the type of motion which affects the accuracy ofalignment. Any activity which causes the aircraft tochange attitude shall be avoided during the alignmentperiod.

c. (CDU) DIM knob Adjust formaximum brightness of CDUdisplays.

d. (CDU) AUTO/MAN/TEST switchAUTO.

e. (CDU) Data selector UL POS orUTM POS.

(Observe coordinates of last present position prior toINS turn-off appear in data displays.)

f. (CDU) AUTO/MAN/TEST switchPress and hold for test. Confirmfollowing on CDU:

(Left and right data displays indicate "88°88.8 N/S"and "88°88.8 E/W" respectively.)

(FROM-TO display indicates "8.8".)

(The following pushbuttons and lamps illuminate:ROLL LIM, HOLD, INSERT/ADVANCE, WYPT CHG,ALERT, BAT (on CDU and MSU), WARN, and READYNAV.)

g. (CDU) AUTO/MAN/TEST switchRelease. Confirm following:

(Data displays indicate coordinates in computermemory; all lamps and pushbuttons illuminated in StepF except INSERT/ADVANCE pushbutton light go out.)

h. If UTM coordinates are to be usedVerify that appropriate Gridcoefficients have been loaded.

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i. Present position data Insert.NOTE

After present position 'has beeninserted and computer hasadvanced to state "7", presentposition cannot be reinsertedwithout downmoding to STBYand restarting alignment.

j. Data selector DSRTK/STS

(Observe that the left data display indicates thedesired track stored in computer memory and that theright data display indicates status "-194.")

NOTEIf the fourth digit from the right, inthe right data display is blank,then a valid heading has not beenstored. Proceed with normalpreflight procedure.

k. Monitor the right data display for achange from alignment state "9" toalignment state "8". This will beindicated by a reading of "--184".

l. Monitor the right data display formalfunction codes (table 3-5).

(Loss of 26 VAC is indicated by an illuminatedWARN lamp and a reading of ".03184" in the right-handdisplay. An inoperative magnetic compass system isindicated by an extinguished WARN lamp and a readingof ".03184" in the right-hand display.)

NOTEIf above displays appear, a storedheading alignment may not bepossible.

m. If there are malfunction codes,proceed to ABNORMALPROCEDURES in this chapter.

n. When alignment state "7" isreached, the INS will advance to theNAV mode.

(Observe that "---05" will appear in the STS display.)

NOTETo achieve best accuracy, enginestart and heavy loading activityshould be delayed until entry intoNAV mode.

NOTEWith the data selector in either theL/L WY PT or UTM WY PTposition, waypoint and TACANstation data may be loaded at anytime after turn-on. The operatorshould occasionally switch thedata selector to DSRTK/STS tomonitor system status indicatorsin the right data display.

(2) (CDU) AUTO/MAN/TEST switch Pressand hold. Confirm:

1. Left and right data displays indicate"88°88.8 N/S" and "88°88.8 E/W"respectively.

2. FROM-TO display is "8.8".

3. Following pushbuttons andlamps illuminate: ROLL LIM, HOLD,INSERT/ADVANCE, WYPT CHG,ALERT, BAT (on CDU and MSU),WARN and READY NAV.

(3) TEST switch Release. (Observe datadisplays indicate coordinates in computer memory; alsoall lamps and pushbuttons lighted in step 7 exceptINSERT/ADVANCE pushbutton extinguish.)

(4) Appropriate grid coefficients Verify loaded(if UTM coordinates are to be used.)

(5) Verify UTM grid coefficients.

1. Data selector UTM WYPT.

2. Keys "5" and "9" Press simultaneously.Observe FROM-TO display is blank. Earth flatness

coefficient appears in left display. The relative earthradius, in meters, appears in right display.)

NOTEThese values are retained fromturn-on to turn-on unless changedby operator.

3. Verify that values correspond tothose required for spheroid beingused.

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NOTEValues for various spheroids are

listed in Table 3-1.

4. If values are correct, return CDU tonormal display mode bymomentarily setting data selector toany position except UTM WYPT. Ifvalues are to be changed, continuewith following steps:

NOTEThe INS geographic position, asread in L/L displays, will not beaffected by any changes in thesecoefficients.

5. Keys "2"or "8" Press to indicatefollowing is the flatness coefficient.

(Observe INSERT/ADVANCE pushbutton lightilluminates.)

6. Load earth flatness coefficient bypressing keyboard keys insequence.

(Example: 2 9 4 9 8 = 29498. Observe numberappears in left display as keys are pressed.)

7. INSERT/ADVANCE pushbuttonPress.

(Observe INSERT/ADVANCE pushbutton light goesout.)

8. Keys "4" or "6" Press, to indicatefollowing load is relative earthradius.


9. Load relative earth radius bypressing keyboard keys insequence.

(Example: 8 2 0 6 m = 8206. Observe numbersappear in the right display as keys are pressed.)

NOTEZone symbol is to be ignored.

Table 3-1. Various Values for UTM Grid Coefficients


(Observe INSERT/ADVANCE pushbutton lightextinguishes.)

11. Data selector UTM POS.(Observe coordinate will reflect values of new

spheroid.)(6) Insert present position.

(7) Perform abbreviated INS interface test Asrequired:

NOTEAssuming a level aircraft, attitudeindicators will become levelduring alignment State "8" andremain level in all modes until INSis shut down. Warning indicatorsfor INS attitude signals from INSare valid while attitude spheredisplay is level.

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NOTEThe INS can provide test signals to theHorizontal Situation Indicator (HSI) andconnected displays. Pressing TEST switchduring Standby, Align, or NAV mode causesall digits on connected digital displays toindicate "8's," and lights the HSI and theALERT lamp. Additional HSI test signals areprovided when INS is in Align and dataselector is at any position other thanDSRTK/STS. Under those conditions,pressing TEST switch causes HSI to indicateheading, drift angle, and track angle error allat "0°" or "30°." At the same time, crosstrack deviation is indicated at "3.75" nauticalmiles (one dot) right or left and INS-controlled HSI flags are retracted from view.

NOTEOutput test signals are supplied to theautopilot when INS steering is selected.Rotating AUTO/MAN switch to AUTO andpressing TEST during Align furnishes a 15°left bank steering command. A 15° rightbank steering command is furnished whenAUTO/MAN switch is set to MAN.

NOTEThe quick test procedure may be performedany time after Alignment State "8" is reachedand prior to entry into NAV.

1. Mode selector ALIGN.(Observe CDU displays are illuminated.)

2. Data selector DSRTK/STS(monitor right data display untilState "8" (or lower) is reached.)(Observe right data display is ---N4,where "N" is not "9".)

3. AUTO/MAN switch MAN.

4. INS Couple to flight director andautopilot, as applicable.

5. Data selector Set to any positionexcept DSRTK/STS.

6. TEST switch Press and hold.

NOTEAfter performing the precedingstep, observe:

MSU All lamps illuminated.CDU All lamps illuminated. All "8's"

displayed.HSI All angles 30°. Cross-track devia-

tion bar is one dot right. All INSflags are retracted.

Flight Director/ A 15° steering command is is-Autopilot sued.Mission Control LINK UPDATE annunciated.Panel TACAN UPDATE annunciated.

7. CDU TEST switch Hold depressed,and rotate AUTO/MAN switch toAUTO.

(Observe all indications are as in step (6) except a15' left steering command is issued. On HSI, All anglesare "0°" and cross-track deviation bar is one dot left.)

8. CDU TEST switch Release. Ifdesired, decouple INS.

(Observe operation returns to normal.)

NOTEPrior to pressing INSERT/ADVANCEpushbutton, any incorrectly loaded data canbe corrected by pressing CLEAR pushbuttonand reloading correct data.

NOTEWhile parked aircraft is undergoingalignment, encoding altimeter will supplythe field elevation (aircraft altitude) into INS.

NOTEOnce present position has been insertedand computer has advanced to alignmentstate "7", present position cannot bereinserted without downmoding to STBYand restarting alignment.

NOTEIf longitude and latitude coordinates arebeing used, skip step 8 (a) and proceed withstep 8 (b).

(a) Insert UTM coordinates of aircraftpresent position:

1. Data selector UTM POS.(Observe that prior to initial load, INSERT/

ADVANCE pushbutton light illuminates.)

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2. To load zone and easting valuesPress keys in sequence, startingwith"E".

(Example: Zone 16, 425 km East = E16 425.Observe that zone and easting in kilometers appear inright data display as keys are pressed.)


(Observe pushbutton light remains illuminated.)4. To load northing data Press keys

in sequence, starting with "N" or "S"to indicate north or southhemisphere.

(Example: 4749 km North = N 4749) Observenorthing kilometers appear in left data displays as keysare pressed.)


(Observe pushbutton light remains illuminated.)6. INSERT/ADVANCE pushbutton

Press.(Observe extra precision display for present position

northing and easting, to the nearest meter, appears inleft and right data displays, respectively.)

7. To load extra precision easting dataPress keys in sequence, startingwith "E".

(Example: 297 m East = E 297) Observe thateasting meters appear in right data display as keys arepressed.)


(Observe pushbutton light remains illuminated.)9. To load extra precision northing

data Press keys in sequence,starting with "N" or "S".

(Example: 901 m North = N 901) Observe thatnorthing meters appear in left data displays as keys arepressed.)

NOTEThe value is always added tonormal value regardless of whichkey (N/S) is pressed to initiate theentry. The normal entryestablishes the hemisphere.


(Observe latitude and longitude data is displayed inUTM and INSERT/ ADVANCE pushbutton lightextinguishes.)

NOTEThe computer will convertcoordinates in the overlap area;however display values willreference appropriate zone.

NOTEThe "W" key may be used toinitiate easting entries; howevercomputer will always interpretsuch entries as an "E" input. "E"will be displayed in normal UTMdisplay.

NOTEExtra precision values are alwaysadded to normal values. As anexample, South 4, 476,995 m willdisplay "4476S" in normal displayand "995" in extra precisiondisplay. There is no roundingbetween the two displays.

(b) To insert geographic coordinates ofaircraft present position:

NOTEPrior to pressingINSERT/ADVANCE pushbutton,any incorrectly loaded data can becorrected by pressing CLEARpushbutton and loading correctdata.

1. Data selector L/L POS.(Observe that, prior to initial load, the INSERT/

ADVANCE pushbutton light is illuminated.)2. To load latitude data Press keys in

sequence, starting with "N" or "S" toindicate north or south.

(Example: 42°54.0' North = N 4 2 5 4 0. Observethat latitude appears in left data display as keys arepressed.)


(Observe pushbutton light remains illuminated.)4. To load longitude data Press keys in

sequence, starting with "W" or "E"to indicate west or east.

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(Example: 87°54.9' West = W 8 7 5 4 9. Observethat longitude appears in right data display as keys arepressed.)

5. INSERT/ADVANCE pushbutton-Press.

(Observe pushbutton light extinguishes.)(8) To insert magnetic variation. (Only if

compass is not operational)

1. Data selector L/L WYPT.

2. Keys "3" and "9" Presssimultaneously.

(Observe that magnetic variation, to a tenth of anarc-minute, appears in the right data display.)

NOTEMagnetic variation is retainedfrom the last NAV mode to beused for reasonableness testsduring stored heading alignment.

NOTEDuring NAV mode, magneticvariation is computed whenaircraft roll is less than 9° andmagnetic heading input is valid(absence of Error Code 17).

3. If magnetic variation is to beloaded, proceed with next two steps.

4. Magnetic variation Load bypressing keyboard keys insequence, starting with "E" or "W"to indicate east or west.

(Example: 14'53.6' East = E 1 4 5 3 6) Observe thatINSERT/ADVANCE pushbutton light illuminates whenfirst key is pressed, and magnetic variation appears inright data display).


(Observe pushbutton light extinguishes.)

NOTEIf in NAV mode and magneticheading input is valid, computerwill reject loaded value.

6. To return INS to normal displaymode, momentarily set dataselector to any position except L/LWYPT.

(9) Data selector DSRTK/STS. Confirm.

1. Left data display indicates desiredtrack angle in computer memory.

2. Right data display indicates ---84, ---74, ---64, or ---54, depending onwhich alignment state the computerhas reached.

(10) To program destinations or TACANcoordinates.

NOTEIf latitude and longitudecoordinates are being used, skipstep I (a) and proceed with step 1(b). Enter all of the data for agiven destination or TACANbefore starting to enter data foranother.

1. To insert waypoint coordinates:

(a) Insertion of UTM waypointcoordinates:

1. Data selector UTM WYPT.Data displays will indicate last coordinates inserted

into related waypoint.2. Thumbwheel Set to waypoint

number to be loaded.

NOTEUTM data may be loaded in anyorder and, until final entry, a valuemay be reloaded.

3. To load zone and easting Presskeys in sequence, starting with "E".

(Example: Zone 16, 425 km East = E16 425.Observe that zone and easting in kilometers appear inthe right data display as keys are pressed.)


(Observe pushbutton light is illuminated.)5. To load northing Press keys in

sequence, starting with "N" or "S"to indicate north or southhemisphere.

(Example: 4749 km North = N 4749. Observe thatnorthing kilometers appear in the left data display askeys are pressed.)

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6. INSERT/ADVANCE push-button - Press.

(Observe pushbutton light remains illuminated.)7. INSERT/ADVANCE push-

button - Press.(Observe that an extra precision display related to

resident value of northing and easting, to the nearestmeter, appears in left and right data displays,respectively.)

8. To load extra precisioneasting value - Press keys insequence starting with "E".

(Example: 297 m East = E 297. Observe thateasting meters appear in the right data display as keysare pressed.)


(Observe pushbutton light remains illuminated.)10. To load extra precision

northing value - Press keys insequence, starting with "N" or"S".

(Example: 901 m North = N 901. Observe thatnorthing meters appear in left data display as keys arepressed. The value is always added to the normal valueregardless of which key (N/S) is pressed to initiate theentry. The normal entry establishes the hemisphere.)


(Within 3 seconds computer converts input intolatitude and longitude for storage in memory. Thestored value is again converted to UTM for display. TheINSERT/ADVANCE pushbutton light extinguishes.Conversion routines may cause displays to change byup to 10 m.)

NOTE

The computer will convertcoordinates in overlap area;however, data display values willreference appropriate zone.

NOTE

The "W" key may be used toinitiate easting entries; however,the computer will always interpretsuch entries as an "E" input. "E"will be displayed in normal UTMdata display.

NOTE

The extra precision values are alwaysadded to normal values. As an

example, South 4,476,995 m willdisplay "4476 S" in the normal displayand "995" in extra precision display. Inother words, there is no roundingbetween the two displays.

12. Repeat steps 2 through 11 foreach waypoint to be loaded.

NOTE

A load cycle may be terminated priorto insertion of all four values bymoving data selector or thumbwheel.

(b) Insertion of geographic waypointcoordinates:

1. Data selector - L/L WYPT.Data displays indicate last coordinates inserted into

the selected waypoint.2. Thumbwheel - Set to

waypoint number to beloaded.

3. To load latitude - Press keysin sequence, starting with "N"or "S" to indicate north orsouth.

(Example: 42°54.0' North = N 4 2 5 4 0. Observethat INSERT/ADVANCE pushbutton light illuminateswhen first key is pressed, and latitude appears in leftdata display as keys are pressed.)


(Observe pushbutton light extinguishes.)5. To load longitude - Press

keyboard keys in sequence,starting with "W" or "E"indicating west or east.

(Example: 87°54.9' West = W 8 7 5 4 9. Observethat INSERT/ADVANCE pushbutton light illuminateswhen first key is pressed, and longitude appears indisplay as keys are pressed.)


(Observe pushbutton light extinguishes.)7. If desire to insert extra

precision coordinate data -Press INSERT/ ADVANCEpushbutton.

(Observe that arc-seconds for loaded latitude andlongitude, to nearest tenth of a second, appear in leftand right data displays, respectively.)

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8. To load related arc-secondvalues for latitude -Press keysin sequence, starting with "N".

(Example: 35.8" North = N 358.)9. INSERT/ADVANCE push-

button- Press.(Observe pushbutton light extinguishes.)

10. To load related arc-secondvalues for longitude -Presskeys in sequence, startingwith "E".

(Example: 20.1" East = E 201.)11. INSERT/ADVANCE push-

button - Press.(Observe pushbutton light extinguishes.)

12. Repeat steps 2 through 11 foreach waypoint to be loaded.

NOTE

In above example, if INSERT/ADVANCE pushbutton was pressed,the following normal display wouldappear: "42°54.5 (N)" and 87°54.3(W).The extra precision values are addedto normal values and normal datadisplays are not rounded off.

NOTE

The normal geographic coordinatesmust always be loaded prior to extraprecision values.

NOTE

The directions "N" or "S" and "E" or"W" are established during normalcoordinate entry. Either key may beused to initiate entry during extraprecision loads and values will beadded to the extra precision valuewithout affecting direction.

NOTE

It is characteristic of the computerdisplay routine to add "0.2" arc-seconds to any display of "59.9" arc-seconds. The value in computer is asloaded by operator.

(11) To insert TACAN coordinates:

(a) Insertion of UTM TACAN stationdata:

NOTE

Prior to pressing theINSERT/ADVANCE pushbutton, anyincorrectly loaded data can becorrected by pressing CLEARpushbutton and loading correct data.

1. Data selector - UTMWYPT.

2. Keys "7" and "9" - Presssimultaneously.

(Observe that number of TACAN station being usedfor navigation flashes on and off in "TO" display anddata displays indicate coordinates of station selected bythumbwheel.)

3. Thumbwheel - Set to numberof station to be loaded.Confirm:

a. Thumbwheel is in de-tent.

b. Station "0" cannot beloaded.

(Observe that if number of station to be loaded issame as number of the TACAN station currently beingused, number in "TO" display will be set to "0" whenTACAN data is loaded.)

c. To load zone and easting- Press keys in sequence,starting with "E".

(Example: Zone 16, 425 km East = E16 425.Observe that zone and easting in kilometers appear inthe right display as keys are pressed.)

d. INSERT/ADVANCEpushbutton - Press.

(Observe that pushbutton light is illuminated.)To load northing - Press keys in sequence, starting

with "N" or "S" to indicate north or south hemisphere.(Example: 4749 km North = N 4749. Observe that

Northing kilometers appear in left data display as keysare pressed.)

e. INSERT/ADVANCEpushbutton - Press.

(Observe pushbutton light remains illuminated.)

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f. INSERT/ADVANCEpushbutton - Press.

(Observe that extra precision display related to theresident value of northing and easting, to nearest meter,appears in left and right data displays, respectively.)

NOTE

UTM data may be loaded in any order.Until final fourth entry, actuation ofINSERT/ADVANCE pushbutton withouta prior data entry will cause normaland extra precision UTM data to bealternately displayed.

g. To load extra precisioneasting value - Presskeys in sequence, startingwith "E".

(Example: 297 m East = E 297. Observe thateasting meters appear in right data display as keys arepressed.)

h. INSERT/ADVANCEpushbutton - Press.

(Observe pushbutton remains illuminated.)

i. To load extra precisionnorthing value - Presskeys in sequence, startingwith "N" or "S".

(Example: 901 m North = N 901. Observe thatnorthing meters appear in left data display as keys arepressed. The value is always added to normal valueregardless of which key (N/S) is pressed to initiate entry.The normal entry establishes hemisphere.)

j. INSERT/ADVANCEpushbutton - Press.

(Observe that during the next 1 to 3 seconds, thecomputer converts input into latitude and longitude forstorage in memory. The stored value is again convertedback to UTM for display to operator. TheINSERT/ADVANCE pushbutton light extinguishes. Theconversion routines may cause data displays to changeby up to 10 m.)

NOTE

The computer will convert coordinatesin overlap area; however, data displayvalues will reference appropriate zone.

NOTE

The "W" key may be used to initiateeasting entries; however, thecomputer will al-

ways interpret such entries as an Einput. "E" will be displayed in normalUTM data display.

NOTE

The extra precision values are alwaysadded to normal values. As anexample, South 4,476.995 m willdisplay "4476 S" in normal display and"995" in extra precision display. Inother words, there is no roundingbetween the two displays.

k. INSERT/ADVANCEpushbutton - Press.

(Observe right data display indicates last previouslyinserted altitude, and left data display is blank.)

l. To indicate the followingload is altitude - Presskeys "4" or "6".


m. To load altitude in feet -Press keys in sequence.

(Example: 1230 ft = 1230. Observe that numbersappear in right data display as keys are pressed.)

NOTE

Altitude inputs are limited to 15,000 feet.

n. INSERT/ADVANCEpushbutton - Press.


o. INSERT/ADVANCEpushbutton - Press.

(Observe that left data display indicates lastpreviously inserted channel number, and right display isblank.)

p. To indicate following loadis channel number -Presskeys "2" or "8"


q. To load channel number -Press keys in sequence.

(Example: 109 = 109. Observe number appears inleft data display as keys are pressed.)

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r. INSERT/ADVANCEpushbutton - press.


NOTE

Any number will be accepted by INS;however, only stations with a channelnumber within range of "1" through"126" will be used for TACAN mixing.

NOTE

Channel number has an implied "X" suffix.

NOTE

Degree symbol (°) should be disregarded whenreading altitude and data display.

s. INSERT/ADVANCEpushbutton - Press.

(Observe station northing, zone, and eastingreappear.)

t. Repeat steps 3 through20 for each TACANstation.

u. To return INS to normalmode, momentarily setdata selector to UTMPOS.

(b) Insertion of geographic TACANstation data:

NOTE

Prior to pressing INSERT/ADVANCEpushbutton, any incorrectly loadeddata can be corrected by pressingCLEAR pushbutton and loadingcorrect data.

1. Data selector - L/L WYPT.

NOTE

If number of station to be loaded issame as number of TACAN stationcurrently being used, number in "TO"display will be set to "O" when TACANdata is loaded.


(Observe that number of TACAN station being usedfor navigation flashes on and off in "TO" display. Data

displays indicate coordinates of station selected viathumbwheel.)

3. Thumbwheel - Set to numberof station being loaded.(Insure thumbwheel is indetent.)

NOTE

Station "0" cannot be loaded.

4. To load latitude - Press keysin sequence, starting with "N"or "S" to indicate north orsouth.

(Example: 42°54.0' North = N 4 2 5 4 0. Observethat INSERT/ADVANCE pushbutton light illuminateswhen first key is pressed.)

5. INSERT/ADVANCE push-button - Press


6. To load longitude - Presskeys in sequence, startingwith "W" or "E" indicatingwest or east.

(Example: 87°54.9' West = W 8 7 5 4 9. Observethat INSERT/ADVANCE pushbutton light illuminateswhen first key is pressed, and longitude appears in datadisplay as keys are pressed.)


(Observe pushbutton light extinguishes.).


(Observe that the arc-seconds related to loadedlatitude and longitude, to nearest tenth of a second,appear in left and right data display, respectively.)

9. If extra precision coordinatedata is to be inserted Presskeys in sequence, startingwith "N", to load related arc-second values for latitude.

(Example: 35.8" North = N 358.)


(Observe pushbutton light extinguished.)

11. To load related arc-secondvalues for longitude -Press

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keys in sequence, starting with "E".

(Example: 20.1" East = E 201.)



NOTE

In above example, ifINSERT/ADVANCE pushbuttonwere pressed, the followingnormal display would appear:"42°54.5 N" and "87°54.3 W". Theextra precision values are addedto normal values and normaldisplays are not rounded off.

NOTE

The normal geographiccoordinates must always beloaded prior to extra precisionvalues.

NOTE

The directions "N" or "S" and "E"or "W" are established duringnormal coordinate entry. Eitherkey may be used to initiate entryduring extra precision loads andthe values will be added to extraprecision value without affectingdirection.

NOTE

It is characteristic of the computerdisplay routine to add 0.2 arc-seconds to any display of 59.9arc-seconds. The value incomputer is as loaded byoperator.

13. INSERT/ADVANCE push-button -Press.

(Observe that right data display indicates lastpreviously inserted altitude, and left data display isblank.)

14. To indicate the following load isaltitude - Press key "4" or "6".


15. To load altitude - Press keys insequence.

(Example: 1230 ft = 1230. Numbers appear in rightdata display as keys are pressed.

NOTE

Altitude inputs are limited to 15,000 feet.

16. INSER/ADVANCE push-button -Press.

(Observe pushbutton light extinguishes.


(Observe that left data display indicates lastpreviously inserted channel number, and right datadisplay is blank.)

18. To indicate the following load ischannel number -Press key "2" or"8".

(Observe INSERT/ADVANCE pushbutton lightilluminates)

19. To load channel number - Presskeys in sequence.

(Example: 109 = 109. Numbers appear in left datadisplay as keys are pressed.)



NOTE

Any number will be accepted bythe INS; however only stationswith a channel number withinrange of 1 through 126 will beused for TACAN mixing.

NOTE

The channel number has an implied

"X" suffix.

NOTE

Decimal points and degreesymbols should be disregardedwhen reading altitude andchannel number displays.


(Observe station latitude and longitude reappear.)

22. Repeat steps 3 through 19 for eachTACAN station.

23. To return INS to normal displaymodes, momentarily set dataselector to L/L POS.

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(12) Designating fly-to-destinations or TACAN.

1. Data selector L/L WYPT orUTM WYPT, as required.

2. Waypoint thumbwheelSelect destination number.

(Observe number of destination waypoint appears in"TO" part of FROM-TO display.)

3. Data selector HDG DA.

(Observe present aircraft heading appears, tonearest tenth of degree, in left data display; also driftangle, to nearest degree, appears in right data display.)

NOTE

Navigation information is nowavailable from INS for display onpilot's RMI and on pilot's andcopilot's HSI's as determined byCOURSE INDICATOR switches.

(13) To fly selected INS course.

1. Pilot's COURSE INDICATORswitch INS.

2. Pilot's RMI select switch INS.

3. Horizontal Situation Indicators(pilot's and/or copilot's HSI)Steer toward indicators.

4. CDU ALERT lamp Monitor.

(Observe illumination approx 1.3 minutes beforereaching point for automatic leg switch. Indicatorflashes on and off after passing a waypoint, ifAUTO/MAN switch is in MAN.)

(14) Aided TACAN operation.

1. Mode selector NAV.

2. Data selector DSRTK/STS.

3. Key "4" Press.

(Observe right data display is "000004" andINSERT/ADVANCE pushbutton light is illuminated.)

4. INSERT/ADVANCEpushbutton Press.

(Observe right data display is "1-XX4" andINSERT/ADVANCE pushbutton light is extinguished.)

NOTE

Every 15 seconds, the INS willselect next eligible TACAN stationin sequence for 3-58 updating. Tobe eligible, TACAN station range

must be between 5 and 150 nmiand channel between 1 and 126.

5. Data selector L/L WYPT orUTM WYPT.


(Observe channel number of the TACAN station beingused for navigation flashes on and off. Data displaysindicate coordinates of station selected via thumbwheel.)

7. To monitor station selectionObserve FROM-TO datadisplay.

(Observe only the number of stations eligible formixing will be displayed. A "0" indicates that none of the9 stations are eligible for selection.)

8. Monitor "TACAN UPDATE"annunciator.

NOTE

Mixing will not be annunciated if:(a) TACAN control isinappropriately set; (b) TACANstation data loaded in error; (c)aircraft look-down angle is greaterthan 30°; (d) horizontal grounddistance is less than two times thealtitude.

9. To return INS display tonormal. Set data selector toany position except WYPT orDIS/TIME.

10. To monitor progress of updateSet data selector to DSRTK/STS. (Observe AccuracyIndex (AI) will decrement to"0".)

NOTE

To insure favorable geometryduring the update process, thefollowing TACAN station criteriashould be observed:

* One station must be at least 15 nmi off course.

* For optimum single TACAN station updating,update should continue until aircraft has passed thestation.

* For optimum dual TACAN updating, use one "off-track" TACAN station and one "on-track" station.

* For optimum multi-TACAN station updating, thestations should be evenly distributed in azimuth aroundthe aircraft.

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11. Waypoint thumbwheel Set tonumber of first TACAN stationto be used.

(15) Switching from aided to unaided inertial.operation.


2. Key "5" Press.

(Observe INSERT/ADVANCE pushbutton lightilluminates; 000005 appears in right data display.)

3 INSERT/ADVANCE pushbuttonPress.

(Observe INSERT/ADVANCE pushbutton lightextinguishes. Data display returns to normal with "5"appearing in first digit of right display.)

NOTE

Benefits of previous aiding aremaintained but no additionalautomatic updates will be made.

(16) To obtain readouts from INS.

NOTE

The computer is assumed to be inthe NAV mode for all data displays.

(a) System status: Data selectorDSRTK/STS.

(Observe numbers indicating system status appear inright data display.)

(b) Geographic present position: Dataselector L/L POS.

(Observe latitude and longitude of present positionappear in left and right data displays, respectively. Bothdisplays are to tenth of a minute.)

(c) UTM position: Data selector UTMPOS.

(Observe northing and zone with easting of presentposition appear in left and right displays, respectively.Both displays are in kilometers.)

(d) True heading: Data selector HDG DA.

(Observe aircraft heading appears in left data displayto nearest tenth of a degree.)

(e) Ground speed: Data selectorTK/GS.

(Observe ground speed appears in right data display tonearest knot.)

(f) Ground track angle: Data selectorTK/GS.

(Observe ground track angle appears in left datadisplay to nearest tenth of a degree.)

(g) Drift angle: Data selector HDG DA.

(Observe drift angle appears in right data display tonearest degree.)

(h) Wind speed and direction: Dataselector WIND.

(Wind direction appears in left data display to nearestdegree and wind speed appears in right display tonearest knot.)

(i) Desired track angle: Data selectorDSRTK/STS.

(Observe desired track angle in left data display tonearest degree.)

(j) Track angle error: Data selectorXTK/TKE.

(Observe track angle error appears in right datadisplay to nearest degree.)

(k) Cross track distance: Data selectorXTK/TKE.

(Observe cross track distance appears in left datadisplay to nearest nautical mile.)

(I) Distance and time to next waypoint:Data selector DIS/TIME.

(Observe distance to next waypoint, shown in "TO"side of FROM-TO display, appears in left data display tonearest nautical mile.)

(Observe time to reach next waypoint at presentground speed appears in right data display to nearesttenth of a minute.)

(m) Extra precision geographic presentposition display:

1. Data selector L/L POS.

(Latitude and longitude of present position, to nearesttenth of a minute, appears in left and right data displays,respectively.)


(Observe arc-seconds related to present positionlatitude and longitude, to nearest tenth of a second,appear in left and right data displays, respectively.)

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(n) Geographic present inertial positiondisplay.


2. HOLD pushbutton Press.

(Observe HOLD pushbutton light illuminates, latitudeand longitude or present inertial position to a tenth ofdegree appear in left and right data displays,respectively.)

NOTE

While HOLD pushbutton light isilluminated, TACAN and data linkupdates are inhibited.


(Observe arc-second related to present inertial positionlatitude and longitude, to nearest tenth of a second,appears in left and right data displays, respectively.)


(Observe coordinates in UTM grid.)


(Observe INS returns to normal operation and HOLDpushbutton light extinguishes.)

(o) UTM present inertial positiondisplay:



(Observe HOLD pushbutton light illuminates.Northing and zone with easting of the present inertialposition in kilometers appear in left and right datadisplays, respectively.)

NOTE



(Observe extra precision values related to presentinertial position northing and easting, to nearest meter,appear in left and right data displays, respectively.)


(Observe coordinates in latitude-longitudecoordinates.)


(Observe INS returns to normal operation and HOLDpushbutton light extinguishes.)

(p) Distance and time to waypoint otherthan next waypoint:

1. WYPT CHG pushbuttonPress.

(Observe WYPT CHG and INSERT/ ADVANCEpushbutton lights illuminate.)

2. Key "0" Press.

(Observe FROM side of FROM-TO data displaychanges to "0".)

3. Key corresponding to desiredwaypoint Press.

(Observe "TO" side of FROM-TO data displaychanges to desired waypoint number.)

NOTE

Do not press INSERT/ADVANCEpushbutton. This would cause animmediate flight plan change.

4. Data selector DIS/TIME.

(Observe distance to desired waypoint appears in leftdata display to nearest nautical mile. Time to reachdesired waypoint at present ground speed appears inright data display to nearest tenth of a minute.)

5. CLEAR pushbutton Press.

(Observe INS returns to normal operation.)

(Observe INSERT/ADVANCE and WYPT CHGpushbutton lights extinguish. Waypoints definingcurrent navigation leg appear in FROM-TO display.)

(q) Distance and time between any twowaypoints:


(Observe WYPT CHG and INSERT/ADVANCEpushbutton lights illuminate.)

2. Keys corresponding to desiredwaypoints Press in sequence.

(Observe desired waypoint numbers appear inFROM-TO data display as keys are pressed.)

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NOTE

Do not press INSERT/ADVANCEpushbutton. This would cause animmediate flight plan change.


(Observe distance between desired waypointsappears in left data display to nearest nautical mile.Time to travel between desired waypoints at presentground speed appears in right data display to nearesttenth of a minute.)

4. CLEAR pushbutton Press.

(Observe INS returns to normal operation.)

(Observe WYPT CHG and INSERT/ ADVANCEpushbutton light extinguishes. Waypoints definingcurrent navigation leg appear in FROMTO data display.)

(r) Distance to any TACAN station:


(Observe distance to next waypoint to nearestnautical mile is in left data display. Time to nextwaypoint to nearest tenth of a minute is in right datadisplay.)


(Observe number of TACAN station being used fornavigation flashes on and off in FROM-TO display.Distance to TACAN station to nearest nautical mile is inleft data display. Time to next waypoint is in right datadisplay.)

3. If in aided TACAN operationMonitor display.

(Observe station number is selected every 15 seconds.)

4. If not in aided TACANoperation Perform steps (5)through (7).


(Observe INSERT/ADVANCE and WYPT CHGpushbutton light illuminates. Station number flashingdiscontinues.)

6. Key indicating desiredTACAN station number -Press.

(Observe number will appear in left digit location ofFROM-TO data display.)

NOTE

If wrong key is pressed, pressCLEAR; displays will revert to thatindicated in step (2).


(Observe INSERT/ADVANCE and WYPT CHGpushbutton lights extinguish. The loaded digit willappear in right position of FROM-TO display and will beflashing on and off. Distance to that station to nearestnautical mile appears in left data display. The rightdisplay continues to display time to next waypoint.)

8. Data selector WIND,momentarily.

(Returns INS to normal display mode.)

NOTE

If in aided TACAN operation and if thedesired station is not being selected,exit aided operation per procedure:"Switching From Aided to UnaidedInertial Operation", perform steps (1)(8), and then return to aided operationper procedure: "Aided TACANOperation".

(s) Coordinates of any waypoint:


2. Waypoint thumbwheel Setdesired waypoint. Observefollowing:

(L/L WYPT: Latitude and longitude of desiredwaypoint, to a tenth of a minute, appear in left and rightdata displays, respectively.)

(UTM WYPT: Northing and zone with easting ofdesired waypoint, to a kilometer, appear in left and rightdata displays, respectively.)

3. INSERT/ADVANCEpushbutton Press. Observefollowing:

(L/L WYPT: The arc-seconds related to desiredwaypoint, latitude, and longitude, to a tenth of an arc-second, appear in left and right data displays,respectively.)

(UTM WYPT: The extra precision display related todesired waypoint northing and easting, in meters,appear in left and right data displays, respectively.)

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NOTE

L/L WYPT: A coordinate is theaddition of values for degrees,whole minutes and seconds.

EXAMPLE: W87° 54' 58.6" = 870 54.9W and 58.6.

UTM WYPT: A coordinate is the addition of thevalues for kilometers and meters.

EXAMPLE: S 2,474,706m = 2474 S and 706.

(t) TACAN station data:



3. Waypoint thumbwheel Setdesired TACAN station.

(Observe number of TACAN station being used fornavigation flashes on and off.)

(L/L WYPT: Latitude and longitude of desiredTACAN station, to tenth of minute, appears in left andright data displays, respectively.)

(UTM WYPT: Northing and zone with easting ofdesired TACAN station, to a kilometer, appear in leftand right data displays, respectively.)

4. INSERT/ADVANCEpushbutton Press. Observefollowing:

(L/L WYPT: The arc-seconds related to desiredTACAN station, to tenth of arc-second, appear in leftand right data displays, respectively.)

(UTM WYPT: The extra precision display related todesired TACAN station northing and easting, in meters,appear in left and right data displays, respectively.)

NOTE

Direction is indicated in normaldata displays.

NOTE

L/L WYPT: A coordinate is theaddition of values for degrees,whole minutes, and seconds.

EXAMPLE: W 87° 54' 58.6" will be displayed as"87°54.9 W" and "58.6".

UTM WYPT: A coordinate is the addition of valuesfor kilometers and meters.

EXAMPLE: S 2,474,706 m will be displayed as "2474S" and "706".


(Observe TACAN station altitude, in feet, will appearin right data display; degree symbol and decimal pointsshould be disregarded. Left data display is blank.)


(Observe TACAN station channel number, in wholenumbers, will appear in left data display; degree symboland decimal point should be disregarded. Right datadisplay is blank.)

NOTE

If INSERT/ADVANCE pushbuttonis pressed, the normalcoordinates indicated in step (3)will be displayed.

NOTE

Waypoint thumbwheel may bemoved at any time and normalcoordinates for new TACANstation will be displayed.

7. Data selector Momentarily toany position other than L/LWYPT, UTM WYPT orDIS/TIME. (Returns INS tonormal operation.)

(u) Magnetic heading.

1. Data selector HDG/DA.

(Observe true heading to nearest tenth of a degreeappears in left data display. Drift angle to nearestdegree appears in right data display.)

2. Keys "3" and "9" Presssimultaneously and hold.

(Observe magnetic heading to nearest tenth of adegree appears in left data display. Drift anglecontinues to be displayed in right data display.)

3. Keys "3" and "9" Release.

(Observe left data display reverts to true heading.)

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(17) INS updating.

(a) Data link updating.

1. Mode selector NAV.


3. Key "3" Press.

(Observe right data display is "000003" andINSERT/ADVANCE pushbutton lamp is illuminated.)


(Observe right data display is "1-XX3" andINSERT/ADVANCE pushbutton light is extinguished.)

5. Monitor "DAT LINK UPDATE"annunciator on missioncontrol panel.

(Observe illumination will be within 30 seconds, ifcorrection of less than 3 + AI nmi is being transmitted toINS.)

6. The Accuracy Index (AI) willdecrement to "O" as updatingprogresses.

(b) Normal geographic present positioncheck and update:

1. Data selector LIL POS.

(Observe latitude and longitude of present positionappear in left and right data displays, respectively.)

2. HOLD pushbutton Press untililluminated.

(Observe latitude and longitude in data displaysfreeze at values present when HOLD pushbutton ispressed.)

NOTE


3. Keys Press in sequence toload latitude of positionreference, starting with "N" or"S" to indicate north or south.

(Example: 42°54.0' north = N 4 2 5 4 0.)

(Observe INSERT/ADVANCE pushbutton lightilluminates when first key is pressed, and latitudeappears in left data display as keys are pressed.)

4. INSERT/ADVANCE pushbutton Press.

(Observe INSERT/ADVANCE pushbutton lightremains illuminated, and previous value of latitudereappears.)

5. Keys Press in sequence toload longitude of positionreference, starting with "W" or"E" to indicate west or east.

(Example: 87°54.9' west = W 8 7 5 4 9. Observelongitude appears in right data display as keys arepressed.)

(Observe INSERT/ADVANCE and HOLD pushbuttonlights remain illuminated. North position error and eastposition error, in tenth of a nautical mile, will appear inleft and right data displays, respectively.)

NOTE

If WARN lamp illuminates,proceed to step 8; otherwiseproceed to step 9.


(Observe action code "02" and malfunction code"49". This indicates that the radial error between theloaded position and the INS position exceeds 38nautical miles. Operator must evaluate possibility thateither INS is in error or reference point position isincorrect. It is possible to force INS to accept updatedposition by setting data selector to L/L POS andproceeding to step (h).

8. If displayed values are withintolerance, press HOLDpushbutton to return INS tonormal operation. If one orboth values are out oftolerance, proceed the step(10).

9. Key "2" Press.

(Observe left data display is "00000 N";INSERT/ADVANCE and HOLD pushbutton lights areilluminated.)

10. INSERT/ADVANCEpushbutton Press

(Observe INSERT/ADVANCE and HOLD pushbuttonlights extinguish. Present position appears in datadisplays. Present position check and update iscomplete.)

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NOTE

Within 30 seconds, computer willprocess correction and revisedpresent position will appear indata display. If AI prior toposition update is 1 or greater,computer will accept over 95percent of correction shown indifference display. If AI is "0",amount of correction acceptedwill be less and is a function oftime in NAV mode and number ofupdates which have been made.

(c) Extra precision geographic presentposition check and update:


2. Key "2" Press.

(Observe INSERT/ADVANCE pushbutton lightilluminates, "000002" appears in right data display.)


(Observe right data display is "1...XX2",INSERT/ADVANCE pushbutton light is extinguished,and any TACAN or data link updating is discontinued.)

4. Data selector L/L POS.

(Observe latitude and longitude of present positionappear in left and right data displays, respectively.)

5. HOLD pushbutton Press(when aircraft passes overknown position reference.)

(Observe HOLD pushbutton light illuminates.Latitude and longitude in data displays freeze at valuespresent when HOLD pushbutton was pressed.)

6. Load latitude by pressing keysin sequence, starting with "N"or "S" to indicate north orsouth.

(Example: 42°54.0' North = N 4 2 5 4 0. Observelatitude appears in left data display as keys arepressed.)

7. INSERT/ADVANCEpushbutton Press

(Observe INSERT/ADVANCE and HOLDpushbuttons remain illuminated.)

8. Load longitude by pressingkeys in sequence, startingwith "W" or "E" indicatingwest or east.

(Example: 87°54.9' West = W 8 7 5 4 9.)


(Observe INSERT/ADVANCE and HOLD pushbuttonlights remain illuminated.)


(Observe arc-seconds related to present positionlatitude and longitude, to nearest tenth of a second, appearin left and right data displays, respectively.)

11. Load related arc-second valuesfor latitude in sequence, startingwith "N".

(Example: 35.8° North = N 358.)


(Observe INSERT/ADVANCE and HOLD pushbuttonsremain illuminated.)

13. Load related arc-second valuesfor longitude in sequence,starting with "E".

(Example: 20.1° East = E 201.)

NOTE

Extra precision values are added tonormal values and normal displaysare not rounded off.

NOTE

Normal latitude-longitudecoordinates must always be loadedprior to extra precision values.

NOTE

Directions "N" or "S" and "E" or "W"are established during normalcoordinate entry. Either key may beused to initiate entry during extraprecision loads and values will beadded to extra precision valueswithout affecting direction.

NOTE

It is characteristic of the computerdisplay routine to add 0.2 arc-seconds to any display of 59.9 arc-seconds. Value in computer isloaded by operator.

14. Proceed to step (6) in procedure:"Normal Geographic

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TM 55-1510-219-10

Present Position Check andUpdate."

(d) UTM present position check andupdate:

NOTE

UTM data may be loaded in anyorder and, until final entry, a valuemay be reloaded.

1. Data selector UTM POS.

(Observe UTM coordinates of present positionappear in data displays.)

2. HOLD pushbutton Press(when aircraft passes overknown position reference.)

(Observe HOLD pushbutton light illuminates.Coordinates in data display freeze at values presentwhen HOLD pushbutton was pressed.)

NOTE


3. Load zone and easting bypressing keys in sequence,starting with "E".

(Example: Zone 16, 425 km East = E16 425.Observe zone and easting in kilometers appear in rightdata display as keys are pressed.)


(Observe INSERT/ADVANCE pushbutton lightremains illuminated.)

5. Load northing by pressingkeys in sequence, startingwith "N" or "S" to indicatenorth or south hemisphere.

(Example: North 4749 km = N 4749. Observenorthing kilometers appear in left data display as keysare pressed.)




(Observe extra precision display related to presentposition northing and easting, to nearest meter, appearsin left and right data displays, respectively.)

8. Load extra precision eastingvalue by pressing keys insequence, starting with "E".

(Example: 297 m East = E 297. Observe eastingmeters appear in right data display as keys arepressed.)

9. INSERT/ADVANCE push-button Press.


10. Load extra precision northingvalue by pressing keys insequence, starting with "N" or"S".

(Example: 901 m North = N 901. ObserveNorthing meters appear in left data display as keys arepressed. The value is always added to normal valueregardless of which key (N/S) is pressed to initiate entry.Normal entry establishes the hemisphere.)

NOTE

The "W" key may be used toinitiate easting entries; however,the computer will always interpretsuch entries as an "E" input.

NOTE

The extra precision values arealways added to normal values.

NOTE

Any data inserted when HOLDpushbutton lamp is notilluminated will be rejected bycomputer.


(Observe INSERT/ADVANCE and HOLD pushbuttonlights remain illuminated. North position error and eastposition error in kilometers will appear in left and rightdata displays, respectively.)

12. If WARN lamp illuminates,proceed to step (13);otherwise proceed to step (9)in procedure: "Extra PrecisionGeographic Present PositionCheck and Update."

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(Observe action code "02" and malfunction code"49". This indicates radial error between loaded positionand INS position exceeds 62 kilometers. Operator mustevaluate possibility that INS is in error or reference pointposition is incorrect. It is possible to force INS to acceptupdated position by setting data selector to UTM POSand proceeding to step (8) of procedure: "ExtraPrecision Geographic Present Position Check andUpdate.")

14. If updating is to be rejectedPress HOLD pushbutton.

(Observe HOLD and INSERT/ADVANCE pushbuttonlights extinguish. INS returns to normal operation.) (e)Position update eradication:

NOTE

This procedure is not consideredcommon. Its use is limited tothose times where an operationalerror has resulted in an erroneousposition fix.


2. Key "1" Press.

(Observe INSERT/ADVANCE pushbutton lightilluminates, 000001 appears in right data display.)


(Observe INSERT/ADVANCE pushbutton lightextinguishes. Within 30 seconds, data displays return tonormal with "0" (normal inertial mode) in last digit ofright display. AI will be set to approximately three timesthe number of hours in NAV.)

(18) Flight course changes.

(a) Manual flight plan change insertion:


(Observe WYPT CHG and INSERT/ ADVANCEpushbuttons illuminate.)

2. Select new FROM and TOwaypoints by pressingcorresponding keys insequence.

(Observe new waypoint numbers appear in FROM-TO data displays as keys are pressed.)

NOTE

Selecting zero as FROM waypointwill cause desired track to bedefined by computed presentposition (inertial present positionplus fixes) and TO waypoint.


(Observe WYPT CHG and INSERT/ ADVANCEpushbuttons extinguish.)

NOTE

Waypoint zero always containsramp coordinates if no manualflight plan changes are made.When a manual flight plan changeis made, present position atinstant of insertion is stored inwaypoint zero.

(19) After landing procedures.

CAUTION

If INS will be unattended for anextended period, it should be shutdown.

CAUTION

Do not leave INS operating unlessaircraft or ground power andcooling air are available tosystem.

NOTE

The INS may be shut down,downmoded to STBY or Alignmode, or operated in thenavigation mode after landing.The determining factor inchoosing course of action isexpected length of time before thenext takeoff.

NOTE

Do not tow or taxi aircraft duringINS alignment. Movement duringalignment requires restartingalignment

.

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(20) Transient stops.

NOTE

Action to be taken during atransient stop depends upon timeavailable and on availability ofaccurate parking coordinates(latitude and longitude.)

(a) Realignment INS operating(recommended if sufficient time and accurate parkingcoordinates are available.)

NOTE

INS can be downmoded to perform arealignment and azimuth gyrocalibration. Alignment to produce analignment state number of "5" can beaccomplished in approximately 17minutes. During the 17 minute period,an automatic azimuth gyrorecalibration is determined on basis ofdifference between inertial presentposition before downmoding andinserted present position. To obtainfurther refinement of azimuth gyro driftrate, calculated on basis of newlycomputed error data, INS can be left inalignment mode for a longer period,allowing the alignment state numberto attain some value lower than "5".

1. Data selector STBY, then toALIGN.

2. Present position coordinatesInsert, according to procedure:"Geographic Present PositionInsertion" or "UTM PresentPosition Insertion."

(b) Realignment INS shutdown. Performcomplete alignment procedures.

(c) Position update (recommended iftime is not available for realignment.)

NOTE

Perform position update using parkingcoordinates in accordance withprocedure: "Insertion of GeographicWaypoint Coordinates." If parkingcoordinates are not available, proceedas follows:

Continue operation in NAV, if INS accuracy appearsacceptable.

Perform position update using best estimate ofparking coordinates.

(d) Downmoding to standby:

NOTE

INS can be downmoded tostandby operation which willmaintain navigation unit atoperating temperature with gyrowheels running. INS isdownmoded to standby asfollows:

1. Mode selector STBY.

Caution

Do not leave INS operating unlessaircraft or ground power andcooling air are available tosystem.

(e) Shutdown:

1. Mode selector OFF.

NOTE

INS will retain inertial presentposition data computed at time INSis downmoded. This value iscompared with present positioninserted for next alignment anddifference is used to determineazimuth gyro drift rate.

d. Abnormal Procedures.

(1) General. INS contains self-testingfeatures which provide one or more warning indicationswhen a failure occurs. The WARN lamp on the CDUprovides a master warning for most malfunctionsoccurring in the navigation unit. Malfunctions in theMSU or CDU will normally be obvious because ofabnormal indications of displays and lamps. A batteryunit malfunction will shut down INS when battery poweris used.

(2) Automatic INS shutdown.

(a) Overtemperature. An overtemp-erature in navigation unit will cause INS to shut down(indicated by blank CDU displays) when mode selectoris at STBY or ALIGN during ground operation. TheWARN lamp on CDU will illuminate and will notextinguish until mode selector is rotated to OFF. Thecooling system should be checked and corrected iffaulty. If cooling system is satisfactory, navigation unitshould be replaced.

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TM 55-1510-219-10Table 3-2. Malfunction Code Check

Table 3-3. Malfunction Indications and Procedures

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TM 55-1510-219-10Table 3-3. Malfunction Indications and Procedures (Continued)

(b) Low battery charge. A low batteryunit charge will cause INS to shut down when INS isoperating on battery unit power. Both WARN lamp onCDU and BAT lamp on MSU will illuminate and notextinguish until the mode selector is set to OFF. Thebattery unit should be replaced when this failure occurs.

(c) Interpretation of failure indications.It is important to be able to correctly interpret failureindications in order to take effective action. Failureindications are listed below under two main categories:WARN light illuminated, and WARN light extinguished.Under each of these categories are listed otherindications which will give the operator sufficientinformation to take action.

1 WARN light on/off

DISPLAY MALFUNCTION/ RECOMMENDEDACTION

Action codes 01, See Table 3-4.02, 03, 04, 05

No action or Indicates computer failure.malfunctioncodes displayed

Improper Indicates NU computer failure.displays

2 WARN light off

DISPLAY MALFUNCTION/ RECOMMENDEDACTION

CDU displays CDU failure indicated.blank, incorrector frozen

NOTE

It is not possible to load displays fromthe keyboard. A temporary failure of anumerical key may prevent dataloading. If a number cannot be loadedinto latitude or longitude displays, afterpressing/ wiggling the key severaltimes, the cause may be the momentaryhang-up of another key. To identify thefaulty key, rotate the data selector toDSRTK/STS. The right digit on rightdisplay will indicate suspect key. Pressand release suspect key several times.To test whether the keyboard problemis corrected, try pressing any othernumerical key. Its number should nowappear as the right digit. If this test issuccessful, press the CLEAR key andreturn data selector to original dataloading position. Otherwise, a CDUfailure is indicated.

(d) CDU BAT indicator is illuminated.

CAUTION

Operation on battery is an indicationthat there may be no aircraft power toblower motor with resultant loss ofcooling. The INS can operate only alimited time (normally 15 minutes) onbattery power before a low voltageshutdown will occur. Then, immediatecorrective action must be taken.

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NOTE

CDU BAT indicator will illuminatefor 12 seconds in alignment State"8" (about 5 minutes after turn-on). This is normal and indicatesa battery test is in progress. Nocorrective action is required.

NOTE

Ground operation on batterypower should not exceed 5minutes.

(e) To determine corrective action:(Monitor CDU displays while rotating the CDU selectorswitch.)

1. If displays are frozen (do notchange while data selector isbeing rotated) problem isnormally in the navigationunit.

2. If displays respond normallyto the data selector, theproblem is normally absenceof 115V AC power to INS.

(f) For corrective action: Check toassure proper settings of following switches and circuitbreakers essential to INS operation:

1 Overhead circuit breaker panel(fig. 2-26): Circuit breakers In:

1. * AVIONICS MASTERCONTR

2. * INS CONTROL3. * AVIONICS MASTER

PWR #14. * AVIONICS MASTER

PWR #2

2 Overhead control panel (fig.2-18): INVERTER No. 1 or INVERTER No.2 switch On(either).

3 Mission control panel (fig. 4-1):

1. 3-phase AC BUS switchOn.

2. INS POWER & AC CONTcircuit breakers - In.

4 Mission AC/DC power cabinet:INS PWR circuit breaker In.

Table 3-4. Action Codes and Recommended Action

NOTE

CDU BAT indicator shouldextinguish after above correctiveaction. If it remains illuminated,INS will eventually shut downwhen battery voltage drops belowapproximately 19VDC. Flight crewshould prepare for shutdown.

(3) Malfunction indications and procedures.Table 3-2 details the procedure for a Malfunction CodeCheck. Table 3-3 lists a number of malfunctionindications which occur under given modes of operation.Follow procedure given. Table 3-4 details action codesand recommended action. Table 3-5 lists failed testsymptoms by malfunction codes and lists codes forrecommended actions.

(4) Read computer memory through CDU(look routine).

1. WYPT CHG key Press.

2. Keys Enter "99". Note FROMTO displays "99".


4. Data selector LIL WYPT.

5. CDU Enter "N" followed byoctal address of desired (evennumbered) memory location.

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Table 3-5. Malfunction Codes

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NOTE

Program prevents entry of anaddress higher than 13777.

6. Insure waypoint wheel is at"0".

7. INSERT/ADVANCEpushbutton Press. (Observeaddress will appear in bothdata displays.)

8. Data selector DST/TIME.(Observe most significant halfof desired data appears in leftdisplay and least significanthalf appears in right display.)

9. To obtain data at next highermemory locations -Advancewaypoint thumbwheel. (Forexample: If address 400 wasentered with thumbwheel at"0", address 402 will beavailable when thumbwheel isset to "1", "404" whenthumbwheel is set to "2", etc.)

3-29. WEATHER RADAR SET (AN/APN-215).

a. Description. The weather radar set (fig. 3-24)provides a visual presentation of the general sky area of

approximately 120° around the nose of the aircraft,extending to a distance of 240 nautical miles. Thepresentation on the screen shows the location ofpotentially dangerous areas, in terms of distance andazimuth with respect to the aircraft. The radar is alsocapable of ground mapping operations. The weatherradar set is protected by a 5ampere RADAR circuitbreaker located on the overhead circuit breaker panel(fig. 2-26).


(1) Weather radar control-indicator switch.

CONTROL FUNCTION

GAIN control Used to adjust radar receiver gain inthe MAP mode only.

STAB OFF Push type on/off switch. Used toswitch control antennastabilization signals.

Range switches Momentary action type switches.When pressed, clears the screenand increases or decreases therange depending on switch pressed.

TILT control Varies the elevation angle of radarantenna a maximum of 15

Figure 3-24. Weather Radar Control-Indicator (AN/APN-215)

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TM 55-1510-219-10

degrees up or down from horizontalattitude of aircraft.

60° switch Push type on/off switch. Whenactivated. reduces antenna scanfrom 120° to 60°.

TRACK switches Momentary action type switches.When activated, a yellow track lineextending from the apex of thedisplay through top range markappears and moves either left orright, depending on the switchpressed. The track line position willbe displayed in degrees in the upperleft corner of the screen. The linewill disappear approximately 15seconds after the switch is released.It will then automatically return to"0" degrees.

HOLD switch Push type on/off switch. Whenactivated, the last image presentedbefore pressing the switch isdisplayed and held. The wordHOLD will flash on and off in theupper left corner of the screen.Pressing the switch again willupdate the display and resumenormal scan operation.

Function switch Controls operation of the radar set.

OFF Turns set off.

STBY Places set in standby mode. Thisposition also initiates a 90 secondwarmup delay when first turned on.

TEST Displays test pattern to check forproper operation of the set. Thetransmitter is disabled during thismode.

ON Places set in normal operation.

MODE switches Momentary action type switches.Pressing and holding either switchwill display an information list ofoperational data on the screen. Thedata heading will be in blue, all dataexcept present data will be inyellow, and present selected datawill show in blue. The threeweather levels will be displayed inred, yellow, and green. If WXAmode has been selected, the redbar will flash on and off. If theswitch is released and immediatelypressed again, the mode will in-

crease or decrease depending onswitch pressed. When either top orbottom mode is reached, theopposite switch must be pressed tofurther change the mode.

NAV switch Interfaces radar indicator with INSnavigational information to showINS waypoint positions,superimposed on the weatherdisplay, relative to aircraft position.

BRT control Used to adjust screen brightness.

c. Normal Operation.

WARNING

Do not operate the weather radarset while personnel orcombustible materials are within18 feet of the antenna reflectorWhen the weather radar set isoperating, high-power radiofrequency energy is emitted fromthe antenna reflector, which canhave harmful effects on thehuman body and can ignitecombustible materials.

CAUTION

Do not operate the weather radarset in a confined space where thenearest metal wall is 50 feet orless from the antenna reflector.Scanning such surfaces maydamage receiver crystals.

(1) Turn-on procedure. Function switchTEST or ON, as required. (Information will appear aftertime delay period has elapsed.)

(2) Initial adjustments operating procedure.

1. BRT control As required.

2. MODE switches Press andrelease as required.

3. RANGE switches Press andrelease as required.

4. TILT control Move up or downto observe targets above orbelow aircraft level. The echodisplay will change in shapeand location only.

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(3) Test procedure.

1. Function switch TEST.

2. RANGE switches Press andrelease as required to obtain80 mile display.


4. Screen Verify proper display.(The test display consists oftwo green bands, two yellowbands, and a red band on a120-degree scan. The wordTEST will be displayed in theupper right corner. Theoperating mode selected bythe MODE switches, eitherMAP, WX, or WXA, will bedisplayed in the lower leftcorner. If WXA has beenselected, the red band in thetest pattern will flash on andoff. The range will bedisplayed in the upper rightcorner beneath the wordTEST and appropriate rangemark distances will appearalong the right edge of thescreen.)

(a) Weather observation operatingprocedure:

1. Function switch ON.

2. MODE switches Press andrelease as required to selectWX.


4. TILT control Adjust untilweather pattern is displayed.Include the areas above andbelow the rainfall areas toobtain a complete display.

5. MODE switches Press andrelease to select WXA. Areasof intense rainfall will appearas flashing red. These areasmust be avoided.

6. TRACK switches Press tomove track line through areaof least weather intensity.Read relative position indegrees in upper left corner ofscreen.

NOTE

Refer to TM 1-230 for weatherobservation, interpretation andapplication.

(b) Ground mapping operatingprocedure: 1. Function switch ON.

2. MODE switches Press andrelease as required to selectMAP.


4. GAIN control As required topresent usable display.

(4) Standby procedure. Function switchSTBY.

d. Shutdown procedure: Function switch OFF.

e. Emergency operation. Not applicable.

3-30. TRANSPONDER SET (APX-100).

a. Description. The transponder system receives,decodes, and responds to interrogations from Air TrafficControl (ATC) radar to allow aircraft identification,altitude reporting, position tracking, and emergencytracking. The system receives a radar frequency of1030 MHz and transmits preset coded reply pulses on aradar frequency of 1090 MHz at a minimum peak powerof 200 watts. The range of the system is limited to line-of-sight.

The transponder system consists of a combinedreceiver/transmitter/ control panel (fig. 3-25) located onthe pedestal extension; a pair of remote switches, oneon each control wheel; and two antennas, located on theunderside and top of the fuselage. The system isprotected by the 3-ampere TRANSPONDER and the 35-ampere AVIONICS MASTER PWR #1 circuit breakerson the overhead circuit breaker panel (fig. 2-26).

b. Transponder Control Panel fig. 3-25).

CONTROL FUNCTION

TEST/GO Illuminates to indicate successfulindicator completion of built-in-test (BIT).

TEST/MON Illuminates to indicate system indicatormalfunction or interrogation by a groundstation.

ANT switch Selects desired antenna for signal input.

TOP Selects upper antenna.

DIV Selects diverse (both) antennas.

BOT Selects lower antenna.

RAD TEST Enables reply to TEST mode in-

OUT switch terrogations from test set.

3-74

TM 55-1510-219-10

Figure 3-25. Transponder Control Panel (AN/APX - 100)

MASTER Selects system operating mode.CONTROLOFF Deactivates system.STBY Activates system warm-up

(standby) mode.NORM Activates normal operating

mode.EMER Transmits emergency reply code.STATUS ANT Illuminates to indicate the BITindicator or MON fault is caused by high

VSWR in antenna.STATUS KIT Illuminates to indicate the BITindicator or MON fault is caused by exter-

nal computer.STATUS ALT Illuminates to indicate BIT orindicator MON fault is caused by to Alti-

tude Digitizer.IDENT MIC Selects source of aircraft identi-switch fication signal.IDENT Activates transmission of identi-

fication (IP) pulse.MIC Enables either control wheel

POS IDENT switch to activatetransmission of ident signalfrom transponder.

MODE 4 reply Illuminates to indicate a replyindicator has been made to a valid Mode

4 interrogation.MODE 4 Selects monitor mode for ModeAUDIO OUT 4 operation.switchAUDIO Enables sound and sight moni-

toring of Mode 4 operation.LIGHT Enables monitoring REPLY in-

dicator for mode 4 operation.OUT Deactivates monitor mode.MODE 3/A code Select desired reply codes forselectors Mode 3/A operation.MODE 1 code Select desired reply codes forselectors Mode 1 operation.MODE 4 TEST Selects test mode of Mode 4OUT switch operation.TEST Activates built-in-test of Mode 4

operation.ON Activates mode 4 operation.OUT Disables Mode 4 operation.MODE 4 code Selects preset Mode 4 code.control

3-75

TM 55-1510-219-10

M-C, M-3A, Select test or reply mode of re-M-2, and M-1 spective codes.switchesTEST Activates self-test of selected

code. Transponder can also re-ply to ground interrogations inthe selected mode during test.

ON Activates normal operation.OUT Deactivates operation of selected

code.MODE 2 code Select desired reply codes forselectors Mode 2 operation. The cover

over mode select switches mustbe slid forward to display the se-lected Mode 2 code.

POS IDENT When pressed, activates tran-pushbutton sponder identification reply.(control wheels,fig. 2-16)

c. Transponder - Normal Operation.

(1) Turn-on procedure. MASTER switch -STBY. Depending on the type of receiver installed, theNO GO indicator may illuminate. Disregard this signal.

(2) Test procedure.NOTE

Make no checks with the masterswitch in EMER, or with M-3/A codes7600 or 7700 without first obtainingauthorization from the interrogatingstation(s).

1. Allow set two minutes towarm up.

2. Select codes assigned for usein modes 1 and 3/A bydepressing and releasing thepushbutton for each switchuntil the desired numberappears in the proper window.

3. Lamp indicators - Operatepress-to-test feature.

4. M-1 switch - Hold in TEST.Observe that no indicatorlamps illuminate.

5. M-1 switch - Return to ON.6. Repeat steps 4 and 5 for the

M-2, M-3/A and M-C modeswitches.

7. MASTER control - NORM.

8. MODE 4 code control - A.Set a code in the externalcomputer.

9. MODE 4 AUDIO OUT switch -OUT.

(3) Modes 1, 2, 3/A, and/or 4 operatingprocedure.

NOTEIf the external security computeris not installed, a NO GO light willilluminate any time the Mode 4switch is moved out of the OFFposition.

1. MASTER control - NORM.2. M-1, M-2, M-3/A, and/or

MODE 4 ON-OUT switches -ON. Actuate only thoseswitches corresponding to therequired codes. Theremaining switches should beleft in the OUT position.

3. MODE 1 code selectors - Set(if applicable).

4. MODE 3/A code selectors -Set (if applicable).

5. MODE 4 code control - Set (ifrequired).

6. MODE 4 REPLY indicator- Monitor to determine whentransponder set is replying toa SIF interrogation.

7. MODE 4 AUDIO OUT switch -Set (as required to monitorMode 4 interrogations andreplies).

8. MODE 4 audio and/orindicator - Listen and/orobserve (for Mode 4interrogations and replies).

9. IDENT-MIC switch - Press toIDENT momentarily.

10. MODE 4 TEST/OUT switch -TEST.

11. Observe that the TEST GOindicator lamp illuminates.

12. MODE 4 TEST/OUT switch -ON.

13. ANT switch - BOT.14. Repeat steps 4, 5, and 6.

Observe that the TEST GOindicator illuminates.

3-76

TM 55-1510-219-10

15. ANT switch - TOP.16. Repeat step 14.17. ANT switch - DIV.18. Repeat step 14.19. When possible, obtain the

cooperation of aninterrogating station toexercise the TEST mode.Execute the following steps:

a. RAD TEST OUT switch -RAD TEST.

b. Obtain verification frominterrogating station that aTEST MODE reply wasreceived.

c. RAD TEST OUT switch -OUT.

(4) Transponder set identification-positionoperating procedure: The transponder set can makeidentification-position replies while operating in codeModes 1, 2, and/or 3/A, in response to ground stationinterrogations. This type of operation is initiated by theoperator as follows:

1. Modes 1, 2, and/or 3/A - On,as required.

2. IDENT-MIC switch - Pressmomentarily to IDENT, whendirected.

NOTE

Holding circuits within thetransponder receiver-transmitterwill transmit identification-positionsignals for 15 to 30 seconds. Thisis normally sufficient time forground control to identify theaircraft's position. During the 15to 30 second period, it is normalprocedure to acknowledge via theaircraft communications set thatidentification position signals arebeing generated.

NOTE

Set any of the M1, M2, M3/A, M-C,or MODE 4 switches to OUT toinhibit transmission of replies inundesired modes.

NOTE

With the IDENT/OUT/MIC switch setto the MIC position, the POS IDENTbut-

ton must be depressed to transmitidentification pulses.(5) Shutdown procedure.

(a) To retain Mode 4 code in externalcomputer during a temporary shutdown:

1. MODE 4 CODE switch -Rotate to HOLD.

2. Wait 15 seconds.3. MASTER control - OFF.4. To zeroize the Mode 4 code

in the external computer turnMODE 4 CODE switch toZERO.

5. MASTER control - OFF.This will automatically zeroizethe external computer unlesscodes have been retained(step 1. above).

3-31. PILOT'S ENCODING ALTIMETER.

a. Description. The encoding altimeter (fig. 326),provides the pilot with an indication of present aircraftaltitude above sea level. It also supplies information tothe transponder for Mode C (altitude reporting) operationand to the aided inertial navigation system. The circuitis protected by the 5-ampere PILOT'S ALT ENCD circuitbreaker on the overhead circuit breaker panel and the 1-ampere F21 fuse in the No. 1 junction box.

b. Controls and Indicators.CONTROL FUNCTION

MILLIBARS Indicates local barometric pres-indicator sure in millibars. Adjusted by use of set knob.IN HG Indicator Indicates local barometric met- ric pressure in inches of mercury. Adjusted by use of set knob.Drum indicator Indicates aircraft altitude in ten- thousands, thousands, and hundreds of feet above sea level.Needle indicator Indicates aircraft altitude in hundreds of feet with subdivi- sions at twenty-foot intervals.CODE OFF flag Presence indicates loss of power(pilot only) to instrument.ALT indicator Not used.Test button When pressed, reading should decrease by 500 feet.

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TM 55-1510-219-10

Figure 3-26. Pilot's Encoding Altimeterc. Normal Operation.

(1) Turn-on procedure. Encoding altimeterwill operate when transponder is operating with M-Cswitch set to center position.

(2) Operating procedure.1. Barometric set knob - Set desired

altimeter setting in IN. HG.window.

2. CODE OFF flag - Check not visible.3. Needle indicator - Check operation.

NOTE

If the altimeter does not readwithin 70 feet of field elevation,when the correct local barometricsetting is used, the altimeterneeds calibration or internalfailure has occurred. An error ofgreater than 70 feet also nullifiesuse of the altimeter for IFR flight.

d. Emergency Operation. Altimeter circuit breaker- Pull (if encoder fault occurs).

3-78

TM 55-1510-219-10CHAPTER 4

MISSION EQUIPMENT

Section I. MISSION AVIONICS

4-1. MISSION AVIONICS OPERATING INSTRUC-TIONS.

Operating instructions for mission avionics equipmentare published in Chapter 3.

4-2. MISSION CONTROL PANEL.

The mission control panel (fig. 4-1), mounted on thecopilot's sidewall, contains three sections.

a. The top section contains the missioncaution/advisory annunciator panel. Refer to Table 4-1.

b. The center section of the mission control panelcontains the instruments used to monitor the missionequipment. These instruments are the DCvolt/ammeter, AC loadmeter, the antenna azimuthindicator and controller.

c. The bottom section of the mission control panelcontains the mission equipment control switches and themission equipment circuit breakers. Refer to Table 4-2.

Table 4-1. Mission Annunciators

MSN OVERTEMP Yellow Mission equipment is over heating.CRYPTO ALERT Yellow Coded messages being received.PWR SPLY FAL/LT Yellow Mission power out of tolerance.CALL Yellow Receiving transmission on VOW.3 O AC OFF Yellow Three phase AC power fault.BATT FEED OFF Yellow Ground fault detected in battery.MISSION POWER Yellow Mission power is off.LINK MODE Yellow WBDL fault in link or contact.RADOME HOT Yellow Radome heat is too high.LINK SYNC Yellow WBDL has synchronization fault.SPCL EQPT OVRD Yellow Mission power switch is in override.DIPLEXER PRESS Yellow Diplexer has lost pressurization.TWTA STANDBY Yellow WBDL is in standby mode.ANT MALF Yellow Boom antenna is out of position.ANT STOWED Green Boom antenna is in horizontal position.ANT OPERATE Green Boom antenna is in vertical position.RADOME HEAT Green Radome heat is on.MISSION AC ON Green Mission AC is on.DAT LINK UPDATE Green INS is updating with information from Data Link.TACAN UPDATE Green INS is updating with information from TACAN.MISSION DC ON Green Mission DC is on.WAVE GUIDE Green Wave guide is pressurized.EXT AC PWR ON Green External AC power is on.EXT DC PWR ON Green External DC power is on.BT00996

4-1

TM 55-1510-219-10

Figure 4-1. Mission Control Panel

4-2

TM 55-1510-219-10Table 4-2. Mission Control Switches

3 ∅ AC BUS Resets and Energizes 3 0 AC BUS.RESETONOFF

MSN CONT Overrides or turns off the mission DC POWER.OVERRIDEAUTOOFF

1 MSN INV Selects 1 or 2 mission inverter or turnsOFF inverters off.2 MSN INV

DATA LINE Turns data link high voltage on, off or standby.HV ONSTBYOFF

ANT SELECT Selects nose or tail data link antennas. InNOSE auto mode, direction and signalAUTO strength of received data linkTAIL determines selection of nose or tail data link antenna.

ANT OVRD Selects automatic rotating boom operation orOPR POS manual selection of antennaAUTO position.ROTATE

AC EXT Turns external AC power on or off.POWEROFF

MISSION AC VOLT Allows monitoring of mission AC circuit.FREQ & LOADBT00997

Section II. AIRCRAFT SURVIVABILITY EQUIPMENT

4-3. M-130 FLARE AND CHAFF DISPENSINGSYSTEM.

a. Description. The M-130 flare and chaffdispensing system provides effective survivalcountermeasures against radar guided weaponssystems and infrared seeking missile threats. Thesystem consists of two dispenser assemblies withpayload module assemblies, a dispenser control panel,a flare dispense switch, two control wheel mounted chaffdispensing switches, an electronic module assembly,and associated wiring. The flare and chaff dispensingsystem is protected by a 5-ampere circuit breaker,placarded M130 POWER, located on the missioncontrol panel (fig. 4-1).

WARNINGRight engine nacelle dispenser isfor chaff only.(1) Dispenser assemblies. Two

interchangeable dispenser assemblies are mounted onthe aircraft. One is located in the aft portion of the rightnacelle and the other is mounted on the right side of thefuselage. On this aircraft the dispenser in the nacellewill be used for chaff only while the dispenser mountedon the fuselage can be used for either flares or chaff.The selector switch (placarded C-F) on the dispensercan be set for either chaff or flares. The unit alsocontains the sensor for the flare detector. The dispenserassembly breech plate has the electrical contact pinswhich fire the impulse cartridges. The unit also containsthe sequencing mechanism.

4-3

TM 55-1510-219-10(2) Payload module assemblies. A

removable payload module assembly is provided foreach dispenser assembly. Each payload module has 30chambers which will accept either flares or chaff. Flaresor chaff are loaded into the rear-end (studded end) ofthe payload module, and secured in place by a retainingplate.

(3) Electronic module assembly (EM). Theelectronic module assembly contains the programmer,the flare detector and a safety switch. The unit islocated behind the pilot's seat.

(a) Flare detector. The flare detector isprovided to insure that a flare is burning when it isejected from the dispenser payload module. If the initialflare fails to ignite, the detector automatically firesanother flare within 75 milliseconds. If the second flarefails to ignite, the detector will fire a third flare. If thethird flare ignition is not detected, the detector will notfire another flare until the system is activated again bypressing the FLARE DISPENSE switch.

(b) Programmer. The programmer isused for the chaff mode only. It has four switches forsetting count and interval of salvo and burst.

(c) Safety switch. The safety switch(with safety pin and yellow flag) prevents firing of chaff

or flares when the safety pin is inserted. The safety pinshall be removed only while the aircraft is in flight orduring test of the system.

(4) Flare dispenser switch. A singlepushbutton switch placarded FLARE DISPENSE,located on the control pedestal, will fire a flare from thedispenser payload module each time it is pressed. If theFLARE DISPENSE switch is held down, it will dispensea flare every 2.3 seconds.

(5) Control wheel mounted chaff dispenseswitches. Two pushbutton switches placarded CHAFFDISP, one located on top left portion of the pilot's controlwheel and the other located on the top right portion ofthe copilot's control wheel, activates the chaffdispensing system when pressed.

(6) Wing mounted safety switch. A wingmounted safety switch (with safety pin and red flag)located on top of the right wing, just aft of the nacelle,prevents the firing of chaff or flares when the pin isinserted. This safety pin shall be inserted while theaircraft is on the ground and removed prior to flight orduring system test.

(7) Dispenser control panel (DCP). Thedispenser control panel (fig. 4-2) is mounted in thecontrol pedestal. Control functions are as follows:

Figure 4-2. Chaff/Flare Dispenser Control Panel4-4

TM 55-1510-219-10CONTROL FUNCTIONARM-SAFE switch When in the SAFE position,

power is removed from theM130 system. When in theARM position, power is appliedto the M-130 system.

ARM light An amber press to test indicatorlight placarded ARM illuminateswhen the ARM-SAFE switch isin the ARM position, when thesafety pins are removed fromthe electronic module and thewing safety switch. Clockwiserotation will dim the indicatorlight.

RIPPLE FIRE switch A guarded switch placardedRIPPLE FIRE fires allremaining flares when moved tothe up position. It is used in theevent of an inflight emergencyto dispense all flares from thedispenser payload module.

FLARE counter Indicates the number of flaresremaining in the dispenserpayload module.

FLARE counter Facilitates setting FLAREsetting knob counter to the number of flares

in the payload module beforeflight.

CHAFF counter Indicates the number of chaffsremaining in the payloadmodule.

CHAFF counter Facilitates setting CHAFFsetting knob counter to the number of chaffs

in the payload module beforeflight. SELECTOR SWITCH

MAN Bypasses the programmer andfires one chaff each time one ofthe chaff dispense switches ispressed.

PRGM Chaff is fired in accordance withthe preset chaff program as setinto the electronic module(count and interval of bursts andsalvo).

(8) Ammunition for dispenser. Ammunitionfor the system consists of countermeasure chaff Ml,countermeasure flares M206, and impulse cartridgesM796.

(a) Countermeasure chaff M1. Theseunits consist of a plastic case 8 inches in length and

0.97 inches square. The base of the chaff case isflanged to provide one-way assembly into the dispenserpayload module. The chaff consists of aluminumcoated fiberglass strands.

(b) Countermeasure flare M206. Theseunits consist of an aluminum case 8 inches in length and0.97 inches square. The base of the flare is flanged toprovide one-way assembly into the payload module.The flare material consists of a magnesium and tefloncomposition. A preformed packing is required in thebase of the flare unit prior to inserting the impulsecartridge.

(c) Impulse cartridge M796. Thiscartridge fits into the base of either the flare or chaff andis electrically initiated to eject flares or chaff from thedispenser payload module.

b. Normal Operation.

NOTEIf aircraft is to be flown with flaredispenser assembly removed,fairing should be removed fromfuselage.

(1) General. At the present time surface to-air intermediate range guided missiles that are launchedagainst the aircraft must be visually detected by theaircraft crew. Crew members must insure visualcoverage over the ground area where a missile attack ispossible. The aircraft radar warning system will onlyalert the pilot and copilot when the aircraft is beingtracked by radar-guided anti-aircraft weapons systems.It will not indicate the firing of weapons against theaircraft.

(2) Crew responsibilities. The pilot ordesignated crew member is responsible for removingthe safety pin from the right wing before flight, and forreplacing it immediately after flight. After the aircraft isairborne the pilot is responsible for removing the safetypin from the electronic module and moving the ARM-SAFE switch on the dispenser control panel to ARM.Before landing, he is responsible for re-inserting thesafety pin in the electronic module and moving theARM-SAFE switch to SAFE. While airborne the pilotand copilot are responsible for scanning the terrain formissile threats. When either pilot recognizes a missilelaunch he will press the FLARE DISPENSE button toeject flares.

(3) Conditions for firing. The dispensersystem should not be fired unless a missile launch isobserved or radar guided weapons systems is detectedand locked on. If a system malfunction is suspected,aircraft commander may authorize attempts to dispenseflares or chaff as a test in a non-hostile area.

4-5

TM 55-1510-219-10

WARNINGAircraft must be in flight to

dispense flares.

(a) Firing procedure.

1. Flares. Upon observing a missilelaunch the designated crewmember will fire a flare. Ifmore than one missile launch is observed, the firingsequence should be continued until the aircraft hascleared the threat area.

2. Chaff. Upon receiving an alertfrom the aircraft radar warning system, the designatedcrewmember will fire the chaff and initiate an evasivemaneuver. The number of burst/salvo and number ofsalvo/program and their intervals as established bytraining doctrine will be set into the programmer prior totake-off (refer to TM 9-1095206-13 for information onsetting programmer). If desired, the operator mayoverride the programmed operational mode and firechaff countermeasures manually by moving thedispenser function selector switch to MANUAL andpressing a dispenser switch.

4-4. SYSTEM DAILY PREFLIGHT/RE-ARM TEST.

The following test procedures shall be conducted prior tothe first flight of each day and prior to each re-arming ofthe dispensers. The first dispenser tested shall be theone used to dispense flares and the second one shall bethe one used to dispense chaff. Notify AVUM if anyimproper indications occur during the tests.

WARNINGAssure payload module is notconnected to dispenser assemblyat any time during the followingtest procedure.

1. On flare dispenser assembly, assure the C-Fselector switch is in F (flare) position.

2. Obtain M-91 test set and ensure that TESTSEQUENCE switch is in START/HOMEposition.

3. Connect base plate of test set to breech ofdispenser assembly. Secure both mountingstuds uniformly hand tight, using 5/32 inch

hexagonal wrench provided in test set carryingcase.

4. Obtain test set power cable from the M-91 testset carrying case and connect cable betweenexterior connection J1 (28V DC) on aircraft andaircraft power + 28V DC (J1) of test set.

5. Remove safety pin from EM and in the top skinof the right wing.

CAUTIONOn DCP, assure that RIPPLE FIREswitch guard is in down position.

6. Provide aircraft power to DCP by setting M130POWER circuit breaker to ON position.

7. On DCP, press ARM lamp. Lamp willilluminate. Release ARM lamp. Lamp willextinguish.

8. On DCP, set FLARE counter to 30 CHAFFCOUNTER to 30 and MAN-PRGM switch toMAN position.

9. On DCP, set ARM-SAFE switch to ARM. ARMlamp will illuminate.

NOTEWhen the test set is installed onthe dispenser assembly and 28volts DC aircraft power has beenapplied, the sequencer switchinside of dispenser assemblyresets, making an audible soundas it rotates. There will be no suchsound if the sequencer switch hasbeen previously reset or if switchis in position 12 or 24.

NOTEOn test set, TS PWR ON lamp(clear) illuminates and remainsilluminated throughout the testsequence until aircraft power totest set (via test set power cable)is disconnected or shut off.

10. Set mission chaff program on EM.11. Perform the following operations on the M-91

test set: a. Press to test the remaining threelamps on test set. Each lamp will illuminate.

4-6

TM 55-1510-219-10NOTE

Replace any lamp that does notilluminate when pressed. If noneof the indicating lamps illuminate,return test set to AVUM.

b. Rotate TEST SEQUENCEswitch clockwise to the nextposition, TS RESET. No visualindication will occur.

c. Rotate TEST SEQUENCEswitch clockwise to SV SELF TESTposition. STRAY VOLTAGE lamp(red) will illuminate.

d. Rotate TEST SEQUENCEswitch clockwise to next position,TS RESET. STRAY VOLTAGElamp (red) will extinguish.

e. Rotate TEST SEQUENCEswitch clockwise to next position,STRAY VOLT. STRAY VOLTAGElamp (red) should not illuminate.

f. Rotate TEST SEQUENCEswitch clockwise to next position,SYS NOT RESET. SYS NOTRESET lamp (amber) should notilluminate. If lamp illuminates,press and release MANUALSYSTEM RESET switch and SYSNOT RESET lamp should thenextinguish.

NOTEWhen the MANUAL SYSTEMRESET switch is pressed andreleased, and 28 volts DC powerhas been applied, the sequencerswitch inside the dispenserassembly resets, making anaudible sound as it rotates. If thesequencer switch has beenpreviously reset or if the switch isin position 12 or 24, there will beno such sound.

g. Rotate TEST SEQUENCE switchclockwise to next position, DISPCOMP.

12. Press FLARE DISP switch once. For eachdepressing, the FLARE counter on DCP shouldcount down in groups of three.

13. On DCP, raise RIPPLE FIRE switch guard andset toggle switch to up position until FLAREcounter counts down to 00. Return switch guardto down position. On DCP, reset FLAREcounter back to 30. DISPENSER COMPLETElamp (green) on test set will illuminate.

14. Perform the following operations on the M-91test set:

a. Rotate TEST SEQUENCEswitch counter-clockwise to SYSNOT RESET position. SYS NOTRESET lamp (amber) willilluminate. DISPENSERCOMPLETE lamp (green) willremain illuminated.

b. Press and release MANUALSYSTEM RESET switch. SYS NOTRESET lamp (amber) willextinguish.


c. Rotate TEST SEQUENCEswitch counterclockwise to STRAYVOLT position. STRAY VOLTAGElamp (red) should not illuminate.

d. Rotate TEST SEQUENCEswitch counterclockwise toSTART/HOME position.

NOTEWhen the TEST SEQUENCEswitch is turned to theSTART/HOME position, theDISPENSER COMPLETE lamp willextinguish, the STRAY VOLTAGElamp will illuminate and then willextinguish when passing throughthe TS RESET position.

15. On CHAFF dispenser assembly, assure that C-Fselector switch is in C (chaff) position.

16. Remove M-91 test set from first dispenserassembly.

17. Connect M-91 test set to breech assembly ofsecond dispenser assembly. Secure bothmounting studs uniformly hand tight using ballhexagonal key screwdriver provided in test setcarrying case.

4-7

TM 55-1510-219-10

NOTEWhen the test set is installed onthe dispenser assembly and 28volts DC aircraft power has beenapplied, the sequence switchinside the dispenser assemblyresets, making an audible soundas it rotates. There will be nosuch sound if the sequencerswitch has been previously resetor if switch is in position 12 or 24.

NOTEOn test set, TS PWR ON lamp(clear) illuminates and remainsilluminated through the testsequence until aircraft power totest set (via test set power cable)is disconnected or shut off.

18. Perform the following operations on the M-91test set: a. Press to test all four lamps on testset. Each lamp will illuminate.

NOTEReplace any lamp that does notilluminate when pressed. If noneof the indicating lamps illuminate,return test set to AVUM.

b. Rotate TEST SEQUENCE switchclockwise to TS RESET position.No visual indication will occur.

c. Rotate TEST SEQUENCE switchclockwise to SV SELF TESTposition. STRAY VOLTAGE lamp(red) will illuminate.

d. Rotate TEST SEQUENCE switchclockwise to next position, TSRESET. STRAY VOLTAGE lamp(red) will extinguish.

e. Rotate TEST SEQUENCE switchclockwise to next position, STRAYVOLT. STRAY VOLTAGE lamp(red) should not illuminate.

f. Rotate TEST SEQUENCE switchclockwise to next position, SYSNOT RESET. SYS NOT RESETlamp (amber) should not illuminate.If lamp illuminates, press andrelease MANUAL SYSTEM RESETswitch and SYS NOT RESET lampshould then extinguish.

NOTEWhen the MANUAL SYSTEMRESET switch is pressed and

released, and 28 volts DC powerhas been applied, the sequencerswitch inside the dispenserassembly resets, making anaudible sound as it rotates. IF thesequencer switch has beenpreviously reset or if the switch isin position 12 or 24, there will beno such sound.

g. Rotate TEST SEQUENCE switchclockwise to next position, DISPCOMPL.

19. Press pilot CHAFF DISP switch once. Presscopilot CHAFF DISP switch once. On DCP, foreach depressing, the CHAFF counter shouldcount down by an increment of one.

20. On DCP, set MAN-PRGM switch to PRGMposition.

21. Press any one of CHAFF DISP switches inaircraft. In DCP, the number shown on CHAFFcounter should decrease in accordance with theprogram set on the EM.

22. Repeatedly press other CHAFF DISPENSEswitch until CHAFF counter on DCP reads 00.

23. On test set, observe DISPENSE COMPLETElamp (green) is illuminated and then perform thefollowing operations:

a. Rotate TEST SEQUENCE switchcounter-clockwise to SYS NOTRESET position. SYS NOT RESETlamp (amber) will illuminate.

b. Press and release MANUALSYSTEM RESET switch. SYS NOTRESET lamp (amber) willextinguish.


c. Rotate TEST SEQUENCE switchcounter-clockwise to STRAY VOLTposition. STRAY VOLTAGE lamp(red) should not illuminate.

4-8

TM 55-1510-219-10

d. Rotate TEST SEQUENCE switchcounter-clockwise to START/HOMEposition.

NOTEWhen the TEST SEQUENCEswitch is turned to the OFFposition, the DISPENSERCOMPLETE lamp will extinguish,the STRAY VOLTAGE lamp willilluminate and then will extinguishwhen the OFF position is reached.

24. Install safety pins.25. Disconnect test set power cable.26. Remove M-91 test set from dispenser assembly

and restore in carrying case along with thepower cable and hexagonal wrench.

27. On DCP, set ARM-SAFE switch to SAFEposition.

28. On DCP, reset CHAFF counter to 30.29. Disconnect aircraft power by pulling the M130

POWER circuit breaker located on the missioncontrol panel (fig. 4-1).

30. Proceed immediately to ammunition loadingprocedures.

4-5. AMMUNITION.

a. Ammunition Loading Procedure.

WARNINGOnly one shipping container is tobe opened at a time. If a shippingcontainer has been opened andonly partially emptied, theremaining contents will besecured in the container with anappropriate type of packagingmaterial or filler to adequatelyprevent jostling. All munitions instorage must be in their originalshipping containers.

1. Place payload module assembly on workbench in approved safe area so that theretaining plate is facing up.

2. Remove retaining plate by unscrewing tworetaining bolts.

3. Insert one flare (or chaff) at a time into eachchamber of payload module.

4. Remove plastic dust cap from each chaff orflare.

CAUTIONPrior to insertion of an impulsecartridge, be sure there ispreformed packing in the flarecartridge. (There will be nopreformed packing in chaffcartridges.) Reinstall anypreformed packing that isinadvertently removed with dustcap. The loading of impulsecartridges into a flare or chaffshall be accomplished one at atime.

5. Insert one impulse cartridge into each flare (orchaff).

6. Install retainer plate assembly by screwing totwo retainer bolts into payload module.

WARNINGThe system must have beentested to assure that there is nostray voltage and all aircraftpower must be removed from thesystem prior to unloading thepayload module.

7. On the dispenser control panel, assure ARM-SAFE switch is in SAFE position.

8. On the electronic module and right wing assuresafety pins and flag assemblies are installed.

9. Slide payload module assembly into dispenserassembly and secure two stud bolts, hand tight,using 5/32 inch hexagonal wrench.

b. Ammunition Unloading Procedure.

WARNINGAll aircraft power to the dispensersystem must be turned off prior toremoval of payload module fromdispenser assembly. Safety pinflag shall be installed in theelectronic module prior to landingand the safety pin flag shall beinstalled in the wing-mountedsafety switch immediately afterlanding.

1. On dispenser control panel, assure ARM-SAFE switch is in SAFE position.

4-9

TM 55-1510-219-10

2. Assure safety pin and flag are inserted intoelectronic module and in the wing mountedsafety switch.

WARNINGIf there is an indication that amisfire occurred, notify EODpersonnel for disposition anddisposal.

3. Remove module from dispenser assembly byunscrewing two stud bolts with a 5/32 inchhexagonal wrench and sliding out of dispenserassembly.

4. Remove retaining plate from payload module byunscrewing two retaining bolts.

5. Remove expended and unexpended impulsecartridges and flares (or chaff) from payloadmodule.

6. Repack unexpended items in original containersand return to stores.

NOTEIt is not unusual for the case of achaff cartridge to crack whenfired. It does not effectperformance of the item andshould not be reported as amalfunction.

4-6. RADAR SIGNAL DETECTING SETS AN/ APR-39(V)1 OR AN/APR-39(V)2.

The radar signal detecting system indicates therelative position of search radar stations. Audio warningsignals are applied to the pilot's and copilot's headsets.The radar signal detecting set is protected by the 7.5-ampere circuit breaker placarded APR-39, located onthe mission control panel (fig. 4-1). The associatedantennas are shown in figure 2-1. For AN/APR-39(V)loperating instructions, refer to TM 11-5841-283-20.Refer to TM 11-584120-2 for the AN/APR-39 (V)2operating instructions. Pattern #1, self test, shall be asshown in figure 4-5.

a. Radar Signal Detecting Set Control PanelFunctions (AN/APR-39(V)1) (fig. 4-3).

CONTROL FUNCTION

PWR switch Turns set On or Off.

SELF TEST Initiates self test.switch

DSCRM switch Turns discriminate function On or Off.

Figure 4-3. Radar Signal Detecting Set Control Panel (AN/APR-39(V)1

4-10

TM 55-1510-219-10

AUDIO control Adjusts audio level.

b. Radar Signal Detecting Set Control PanelFunctions (AN/APR-39(V)2) (fig. 4-4).

CONTROL FUNCTIONPWR switch Turns set On or OffALT/HI Low Selects mode of operationswitchHI Selects high altitude threat

modeLOW Selects low altitude threat modeTEST switch Initiates self testAUDIO control Adjust audio level

c. Radar Signal Detecting Set Indicator Functions(fig. 4-5).

CONTROL FUNCTIONMA indicator Illuminates to indicate the

presence of an MA threat.

BRIL control Adjusts brilliance.

DAY-NIGHT Rotation adjusts intensity ofcontrol display.

4-7. RADAR WARNING RECEIVER AN/APR-44()(V)3.The radar warning receiver (fig. 4-6) indicates thepresence of certain types of search radar signals.The radar warning receiver is protected by the 5amperecircuit breaker placarded APR-44, located on themission control panel (fig. 4-1). For operatinginstructions, refer to TM 11-5841-291-12.

CONTROL FUNCTIONRadar warning Illuminates to indicate theindicator presence of an AI or SAM

threat.

VOLUME Adjusts volume.controlPOWER switch Turns set On or Off.

Figure 4-4. Radar Signal Detecting Set Control Panel (AN/APR-39(V)2

4-11

TM 55-1510-219-10

Figure 4-5. Radar Signal Detecting Set Indicator AP005715

Figure 4-6. Radar Warning Receiver Control Panel AN/APR-44( )(V)3

4-12

TM 55-1510-219-10CHAPTER 5

OPERATING LIMITS AND RESTRICTIONS

Section I. GENERAL5-1. PURPOSE.

This chapter identifies or refers to all operatinglimits and restrictions that shall be observed duringground and flight operations.

5-2. GENERAL.

The operating limitations set forth in this chapterare the direct result of design analysis, tests, andoperating experiences. Compliance with these limits willallow the pilot to safely perform the assigned missionsand to derive maximum utility from the aircraft. Limitsconcerning maneuvers, weight, and center of gravity arealso covered in this chapter.

5-3. EXCEEDING OPERATIONAL LIMITS.

Anytime an operational limit is exceeded anappropriate entry shall be made on DA Form 2408-13.Entry shall state what limit or limits were exceeded,range, time beyond limits, and any additional data thatwould aid maintenance personnel in the maintenanceaction that may be required.

5-4. MINIMUM CREW REQUIREMENTS.

The minimum crew required for flight is two pilots.Additional crewmembers will be added as required, atthe discretion of the commander, in accordance withpertinent Department of the Army regulations.MACOMS may authorize maintenance test flights to beconducted with one qualified pilot at the pilot's station,and a trained technical observer at the copilot's station,in day/visual meteorological conditions only.

Section II. SYSTEM LIMITS

5-5. INSTRUMENT MARKINGS.

Instruments which display operating limitationsare illustrated in figure 5-1. The operating limitationsare color coded on the instrument faces. Color codingof each instrument is explained in the illustration.

5-6. INSTRUMENT MARKING COLOR CODES.

Operating limitations and ranges are illustrated bythe colored markings which appear on the dial faces ofengine, flight, and utility system instruments. Redmarkings indicate the limit above or below whichcontinued operation is likely to cause damage or shortenlife. The green markings indicate the safe or normalrange of operation. The yellow markings indicate therange when special attention should be given to theoperation covered by the instrument. Operation ispermissible in the yellow range, but should be avoided.White markings on the airspeed indicator denotes flapoperating range.

The blue marking on the airspeed indicator denotes bestrate of climb with one engine inoperative, at maximumgross weight, maximum forward c.g., sea level standardday conditions.

5-7. PROPELLER LIMITATIONS.

The maximum propeller overspeed limit is 2200RPM. Propeller speeds above 2000 RPM indicatefailure of the primary governor. Propeller speeds above2080 RPM indicate failure of both primary andsecondary governors. Torque is limited to 81% forsustained operation above 2000 RPM.

5-8. STARTER LIMITATIONS.

The starters in this aircraft are limited to anoperating period of 30 seconds ON, then 5 minutesOFF, for two starter operations. After two starteroperations the starter shall be operated for 30 secondsON, then 30 minutes OFF.

5-1

TM 55-1510-219-10

Figure 5-1. Instrument Markings (Sheet 1 of 3)

5-2

TM 55-1510-219-10


5-3

TM 55-1510-219-10


5-4

TM 55-1510-219-10

5-9. AUTOPILOT LIMITATIONS.a. An autopilot preflight check must be conducted

and found satisfactory prior to each flight on which theautopilot is to be used.

b. A pilot must be seated at the controls with theseat belt fastened when the autopilot is in operation.

c. Operation of the autopilot and yaw damper isprohibited during takeoff and landing, and below 200feet above terrain. Maximum speed for autopilotoperation is 247 knots/0.47 Mach.

d. During a coupled ILS approach do not operatethe propellers in the 1750 to 1850 RPM range.

5-10. FUEL SYSTEM LIMITS.

NOTEAviation gasoline (AVGAS)contains a form of lead which hasan accumulative adverse effect ongas turbine engines. The lowestoctane AVGAS available (less leadcontent) should be used. If anyAVGAS is used the total operatingtime must be entered on DA Form2408-13.

a. Operating Limits. Operation with FUEL PRESSlight on is limited to 10 hours. Log FUEL PRESS lighton time on DA Form 2408-13. One standby boost pumpmay be inoperative for takeoff. (Crossfeed fuel will notbe available from the side with the inoperative standbyboost pump.) Operation on aviation gasoline is timelimited to 150 hours between engine overhaul andaltitude limited to 20,000 feet with one standby boostpump inoperative. Crossfeed capability is required forclimb, when using aviation gasoline above 20,000 feet.

b. Fuel Management. Auxiliary tanks will not befilled for flight unless the main tanks are full. Maximumallowable fuel imbalance is 1000 lbs. Do not take off iffuel quantity gages indicate in yellow arc (less than 265lbs. of fuel in each main tank). Crossfeed only duringsingle engine operation.

c. Fuel System Anti-Icing. Icing inhibitorconforming to MIL-I-27686 PRIST will be added tocommercial fuel, not containing an icing inhibitor, duringfueling operations, regardless of ambient temperatures.The additive provides anti-icing protection and alsofunctions as a biocide to kill microbial growth in aircraftfuel systems.

5-11. BRAKE DEICE LIMITATIONS.

The following limitations apply to the brake deicesystem:

a. The brake deice system shall not be operated atambient temperatures above + 15°C.

b. The brake deice system shall not be operatedlonger than 10 minutes (one timer cycle) with thelanding gear retracted. If operation does notautomatically terminate approximately 10 minutes aftergear retraction, turn the brake deice switch OFF.

c. Maintain 85% NI or higher during simultaneousoperation of the brake deice and surface deice systems.If adequate pneumatic pressure cannot be provided forsimultaneous operation of the brake deice and surfacedeice systems, turn OFF the brake deice system.

d. In order to maintain an adequate supply ofsystems pneumatic bleed air, the brake deice systemshall be turned OFF during single engine operation.

5-12. PITOT HEAT LIMITATIONS.

Pitot heat should not be used for more than 15minutes while the aircraft is on the ground.

5-13. PNEUMATIC SURFACE DEICE SYSTEMLIMITATIONS.

The pneumatic surface deice system shall not beoperated at ambient temperatures below -40'C.

5-5

TM 55-1510-219-10

Section III. POWER LIMITS5-14. ENGINE LIMITATIONS.

Operation of the engines is monitored byinstruments, with the operating limits marked on theface of each instrument.

CAUTIONEngine operation using only theengine driven fuel pump withoutboost pump fuel pressure islimited to 10 cumulative hours.All time in this category shall beentered on DA Form 2408-13 forthe attention of maintenancepersonnel.

CAUTIONUse of aviation gasoline is time-limited to 150 hours of operationduring any Time Between-Overhaul (TBO) period. It may be

used in any quantity with primaryor alternate fuel.

5-15. OVERTEMPERATURE AND OVERSPEEDLIMITATIONS.

a. Whenever the limiting temperatures areexceeded and cannot be controlled by retarding thepower levers, the engine will be shut down and a landingmade as soon as possible.

b. During engine operation the temperatures,speeds and time limits listed in the Engine OperatingLimitations chart (table 5-1) must be observed. Whenthese limits are exceeded, the incident will be enteredas an engine discrepancy in the appropriatemaintenance forms. It is particularly important to recordthe amount and duration of over-temperature and/oroverspeed.

c. Continuous engine operation above 725 °C willreduce engine life.

Table 5-1. Engine Operating Limitations

5-6

TM 55-1510-219-10

5-16. POWER DEFINITIONS FOR ENGINE OPERATIONS.

The following definitions describe the enginepower ratings.

a. Takeoff Power. The maximum power available,limited to periods of five minutes duration.

b. Maximum Continuous Power. The highestpower rating not limited by time. Use of this rating isintended for emergency situations at the discretion ofthe pilot.

5-17. GENERATOR LIMITS.

Maximum generator load is limited to 100% forflight and variable during ground operations. Observethe limits shown in Table 5-2 during ground operation.

Section IV. LOADING LIMITS

5-18. CENTER OF GRAVITY LIMITATIONS.

Center of gravity limits and instructions forcomputation of the center of gravity are contained inChapter 6. The center of gravity range will remainwithin limits, providing the aircraft loading isaccomplished according to instructions in Chapter 6.

5-19. WEIGHT LIMITATIONS.

WARNINGThe ability to sustain loss ofengine power and successfullystop, continue the takeoff, orclimb before or after gearretraction is not assured for allconditions. Thorough mission

planning must be accomplished,prior to takeoff, by analysis ofmaximum takeoff weight,accelerate-stop distance, positiveengine-out climb at liftoff,accelerate-go distance, takeoffclimb gradient, and climbperformance.

The maximum designed gross weight is 14,200pounds for takeoff. Maximum landing weight is 13, 500pounds. Maximum ramp weight is 14,290 pounds.Maximum zero fuel weight is 11,500 pounds.

Section V. AIRSPEED LIMITS MAXIMUM AND MINIMUM

5-20. AIRSPEED LIMITATIONS.

Airspeed indicator readings contained inprocedures, text, and illustrations throughout thisOperator's Manual are given as indicated airspeed(IAS). Airspeed indicator markings (fig. 5-1) andplacarded airspeeds, located on the cockpit overheadcontrol panel (fig. 2-18), are calibrated airspeed (CAS).Airspeed Calibration Charts are provided in Chapter 7.

5-21. MAXIMUM ALLOWABLE AIRSPEED.

Refer to figure 5-2 to determine limiting airspeedsat maximum gross weight under various conditions.The maximum allowable airspeed is 247 KIAS/0.47Mach.

5-22. LANDING GEAR EXTENSION SPEED.

The airspeed limit for extending the landing gearand for flight with the landing gear extended is 184KIAS.

5-23. LANDING GEAR RETRACTION SPEED.

The airspeed limit for retracting the landing gear is166 KIAS.

5-24. WING FLAP EXTENSION SPEEDS.

The airspeed limit for APPROACH extension(40%) of the wing flaps is 202 KIAS. The airspeed

5-7

Table 5-2. Generator Limits

TM 55-1510-219-10

Figure 5-2. Flight Envelope

5-8

TM 55-1510-219-10

limit for full DOWN extension (100%) of the wing flaps is control airspeed (Vmc) at sea level standard conditions is 89157 KIAS. If wing flaps are extended above these speeds, K I A S .the flaps or their operating mechanisms may be damaged.

5-26. MAXIMUM DESIGN MANEUVERING SPEED.5-25. MINIMUM SINGLE-ENGINE CONTROLAIRSPEED (VMC). The maximum design maneuvering speed is 173 KIAS.

Chapter 7, Section X describes minimum single-engine control airspeed. The minimum single-engine

Section VI. MANEUVERING LIMITS

5-27. MANEUVERS. G's with wing flaps up or a positive load factor of2.0 G’s, or a negative 1.227 G’s with wing flaps

a. The following maneuvers are prohibited.down.

1. Spins.b. Recommended turbulent air penetration airspeed is

173 KIAS.

2. Aerobatics of any kind.

3. Abrupt maneuvers above 173 KIAS.

5-28. BANK AND PITCH LIMITS.

a. Bank limits are 60° left or right.

4. Any maneuver which results in a positive load fac-tor of 3.10 G’s or a negative load factor of 1.227

b. Pitch limits are 30° above or below the horizon.

Change 3 5-8.1

TM 55-1510-219-10

SECTION VII. ENVIRONMENTAL RESTRICTIONS

5-29. ALTITUDE LIMITATIONS.

The maximum altitude that the aircraft may be operatedat is 31,000 feet. When operating with inoperative yawdamp, the altitude limit is 17,000 feet.

5-30. TEMPERATURE LIMITS.

a The aircraft shall not be operated when the ambienttemperatures are warmer than ISA +37°C at SL to 25,000feet, or ISA +31°C above 25,000 feet.

b. Engine ice vanes shall be retracted at +15°C andabove.

5-31. FLIGHT UNDER IMC (INSTRUMENTMETEOROLOGICAL CONDITIONS).

This aircraft is qualified for operational in instrumentmeteorological conditions.

5-31A. ICING LIMITATIONS (TYPICAL).

W A R N I N G

While in icing conditions, if there is anunexpected 30% increase of torque neededto maintain airspeed in level flight, acumulative total of two or more inches of iceaccumulation on the wing, an unexplaineddecrease of 15 knots IAS, or an unexplaineddeviation between pilot’s and copilot’sairspeed indicators, the king environmentshould be exited as soon as practicable. Iceaccumulation on the pitot tube assembliescould cause a complete loss of airspeedindication.

The following conditions indicate a possibleaccumulation of ice on the pitot tube assemblies andunprotected airplane surfaces. If any of these conditionsare observed, the icing environment should be exited assoon as practicable.

1. Total ice accumulation of two inches or more on thewing surfaces. Determination of ice thickness can beaccomplished by summing the estimated ice thickness onthe wing prior to each pneumatic boot deice cycle (e.g. fourcycles of minimum recommended ½-inch accumulation.

2. A 30 percent increase in torque per engine requiredto minimum an desired airspeed in level flight (not toexceed 85 percent torque) when operating at recommendedholding speed.

3. A decrease in indicated airspeed of 15 knots afterentering the icing condition (not slower than 1.4 power offstall speed) if maintaining original power setting in levelflight. This can be determined by comparing pre-icingcondition entry speed to the indicated speed after a surfaceand antenna deice cycle is completed.

4. Any variations from normal indicated airspeedbetween the pilot’s and copilot’s airspeed indicators.

5-8.2 Change 3

TM 55-1510-219-10

5-31B. ICING LIMITATIONS (SEVERE).

W A R N I N G

Severe icing may result from environmentalconditions outside of those for which theairplane is certificated. Flight in freezingrain, freezing drizzle, or mixed icingconditions (supercooled liquid water and icecrystals) may result is a build-up onprotective surfaces exceeding the capability ofthe ice protection system, or may result in iceforming aft of these protected surfaces. Thisice may not shed using ice protection systems,and may seriously degrade the performanceand controllability of the airplane.

a. During flight, severe icing conditions that exceedthose for which the airplane is certificated shall bedetermined by the following visual cues. If one or more ofthese visual cues exists, immediately request priorityhandling from air traffic control to facilitate a route or analtitude change to exit the icing conditions:

(1) Unusually extensive ice accreted on theairframe in areas not normally observed to collect ice.

(2) Accumulation of ice on the upper (or lower, asappropriate) surface of the wing aft of the protected area.

(3) Accumulation of ice on the propeller spinnerfarther aft than normally observed.

b. Since the autopilot may mask tactile cues thatindicate adverse changes in handling characteristics, use ofthe autopilot is prohibited when any of the visual cuesspecified above exist, or when unusual lateral trimrequirements or autopilot trim warnings are encounteredwhile the airplane is in icing conditions.

NOTE

All icing detection lights must be operative priorto flight into icing conditions at night. Thissupersedes any relief provided by the masterminimum equipment list (MMEL) or equivalent.

5-32. WIND LIMITATIONS.

The maximum demonstrated crosswind for landing is25 KTS. Wind limitations are described in Chapter 7.

5-33. OXYGEN REQUIREMENTS.

For oxygen requirements, see AR 95-1.

5-34. CABIN PRESSURE LIMITS.

Maximum cabin differential pressure is 6.1 PSI. If acrack in either the cabin window or windshield shouldoccur, refer to chapter 9.

Change 3 5-9

TM 55-1510-219-10

Section VIII. OTHER LIMITATIONS

5-35. INSTRUMENT LANDING SYSTEM LIMITS.

During ILS approach do not operate the propel-lers in the 1750 to 1850 RPM range.

5-36. FERRY CHAIR.

A ferry chair may be installed in the cabin areafor use on ferry missions. The seat may be installedin the forward or the aft facing positions. The side facinglavatory is limited to 170 pounds.

5-37. INTENTIONAL ENGINE CUT SPEED.

Inflight engine cuts below the safe one-engine in-operative speed (Vsse - 104 KIAS) are prohibited.

5-38. RADOME ANTI-ICE OPERATION.

The following limitations apply to operation ofthe radome anti-ice system:

N O T E

Ice accumulation on the forward data linkrandome does not adversely affect the fly-

ing qualities of the aircraft, however iceaccumulation does affect the data link op-eration. Therefore radome anti-ice shouldbe used only in conjunction with the mis-sion equipment.

1. Radome anti-ice shall be off during takeoff,landing, and any single engine operations.

2. Maintain 85% N1 or above during simulta-neous opera t ion o f the radome an t i - i ce ,brake deice, and surface deice systems. If ad-equate pneumatic pressure cannot be pro-vided for simultaneous operation of thesesystems, turn off the radome anti-ice andbrake deice systems.

5-39. CABIN DOOR.

The cabin door is weight limited to 300 poundsto prevent possible structural damage.

5-40. MAXIMUM DESIGN SINK RATE.

The maximum design sink rate below 13,500pounds gross weight is 600 feet per minute. Themaximum design sink rate above 13,500 poundsgross weight is 500 feet per minute.

Section IX. REQUIRED EQUIPMENT FOR VARIOUS CONDITIONS OF FLIGHT

5-41. REQUIRED EQUIPMENT LISTING

a. A Required Equipment for Various Condi-tions of Right (table 5-3) is provided to enable thepilot to identify those systems/components requiredfor flight. For the sake of brevity, the listing doesnot include obviously required items such as wings,rudder, flaps, engines, landing gear, etc. The list alsodoes not include items which do not affect the air-worthiness of the aircraft such as galley equipment,en te r ta inment sys tems , passenger conven ienceitems, etc. It is, however, important to note the ALLITEMS WHICH ARE RELATED TO THAT AIR-WORTHINESS OF THE AIRCRAFT AND NOTINCLUDED ON THE LIST ARE AUTOMATI-CALLY REQUIRED TO BE OPERATIVE

b. It is the final responsibility of the pilot todetermine whether the lack or inoperative status ofa piece of equipment on the aircraft will limit theconditions under which the aircraft may be operat-ed.

(-)Indicates i tem may be inoperative for thespecified Right condition.

(*)Refers to remarks and/or exceptions columnfor explicit information or reference.

Numbered items contained in ( ) are requiredfor flights by AR 95-1.

c. The pilot is responsible for exercising thenecessary operational control to assure that no air-craft is flown with multiple items inoperative, with-out first determining that any interface or interrela-tionship between inoperative systems or componentswill not result in a degradation in the level of safetyand/or cause an undue increase in crew workload.

d. The exposure to additional failures duringcontinued operation with inoperative systems orcomponents must also be considered in determiningthat an acceptable level of safety is being main-tained. The REL may not deviate from requirementsof the operators manual limitations section emer-gency procedures or safety of flight messages.

5-10 Change 2

TM 55-1510-219-10Table 5-3. Required Equipment Listing

Number of items installedSYSTEM VFR Dayand/or VFR Night

COMPONENT IFR DayIFR Night

Icing ConditionsREMARKS and/or Exceptions

AIR CONDITIONINGBleed Air Fail Light 2 - - 1 1 2 Provided bleed air is not used from side of failed

light.Pressurization Controller 1 1 1 1 1 1Safety Valve 1 1 1 1 1 1Outflow Valve 1 1 1 1 1 1Altitude Warning 1 1 1 1 1 1 May be inoperative provided airplane remains

unpressurized.Cabin Rate of Climb 1 1 1 1 1 1Differential Pressure/Cabin Altitude 1 1 1 1 1 1Pressurization Air Source 2 1 1 1 1 1Duct Overtemp Light 1 - - - - - May be inoperative provided bleed air is not used.COMMUNICATIONSInterphone System 1 - - - - -VHF Communications Systems 2 - - - - -Static Discharge Wicks 24 - - 24 24 24 Minimum required - one wick at the outboard end

of each control surface plus top of verticalstabilizer

ELECTRICAL POWERBattery 1 1 1 1 1 1Battery Charge Light 1 1 1 1 1 1DC Generator 2 1 1 2 2 2DC Loadmeter 2 2 2 2 2 2 One may be inoperative provided corresponding

generator caution light Is monitored. One may beDC Generator Caution Light 2 2 2 2 2 2 inoperative provided corresponding loadmeter is

monitored.Inverter 2 1 1 2 2 2Inverter Warning Light 1 - - 1 1 1 May be inoperative provided both inverters

are operative.AC Frequency/Voltmeter 2 2 2 2 2 2EQUIPMENT FURNISHINGSSeat Belts and Shoulder Harnesses;Pilot and Co-Pilot 2 * * * * * * One per installed seat.Emergency Locator transmitter 1 - - - - -

5-11





FIRE PROTECTIONFire Detector System 2 2 2 2 2 2Engine Fire Extinguisher 2 2 2 2 2 2Portable Fire Extinguisher 2 2 2 2 2 2

FLIGHT CONTROLSTrim Tab Indicators - Rudder, Aileron, 3 3 3 3 3 3 May be inoperative provided that theand Elevator tabs are visually checked in the neutral

position prior to each takeoff andchecked for full range of operation.

Flap Position Indicator 1 1 1 1 1 1 May be inoperative provided that theflap travel is visually inspected prior totakeoff.

Flap System 1 - - - - -Rudder Boost 1 - - - - -Yaw Damp 1 1 1 1 1 1 May be inoperative for flight at and be

low 17,000 feet.Stall Warning 1 1 1 1 1 1Autopilot 1 - - - - -

FUEL EQUIPMENTStandby Fuel Boost Pump 2 1 1 1 1 1 Both required for operation on aviation

gasoline above 20,000 feet.Engine Driven Boost Pump 2 2 2 2 2 2Firewall Shutoff Valve 2 2 2 2 2 2Fuel Quantity Indicator 2 2 2 2 2 2 One may be inoperative provided other

side is operational and amount of fuelon board can be established to be adequatefor intended flight. Fuel flow onaffected side must be operational andmonitored.

Crossfeed Valve 1 - - - - - Required for (1) operation with aviationgasoline above 20,000 feet; 12) whenoperating with aviation kerosene whenone standby boost pump is Inoperative.If takeoff with inoperative crossfeed isplanned, mission should be limited tothat range attainable with single engineoperation, one engine supplying fuel.

Crossfeed Light 1 1 1 1 1 1 May be inoperative provided proper operationof crossfeed system is checkedprior to takeoff. Both fuel pressurelights must be operative.

5-12





Fuel Flow Indicator 2 2 2 2 2 2 One may be inoperative provided fuelquantity gages are operative.

Fuel Pressure Warning Light 2 2 2 2 2 2 One may be inoperative providedstandby boost pump operation is ascertainedusing opposite light with crossfeed prior to engine start. Standbyboost pump on side of failed light mustbe operated in flight to assure fuelpressure, should the engine drivenboost pump fail.

Motive Flow Valve 2 - - - - - Required if aux tanks contain fuel.Jet Transfer Pump 2 - - - - - Required if aux tanks contains fuel.Fuel Quantity Gage Selector Switch 1 1 1 1 1 1 May be inoperative provided MAIN

quantity indicators are operational.ICE AND RAIN PROTECTIONAirfoil Deice System (Wing and Horizon- 1 - - - - 1tal Stabilizer)Antenna Deice System 1 - - - - 1Engine Inertial Ice Vanes 2 2 2 2 2 2Ice Vane Lights 4 4 4 4 4 4 May be inoperative provided manual ice

vane controls are operational and used.Windshield Heat, Left and Right 2 - - - - 1 Right side may be inoperative.Windshield Wiper 2 - - - - -Auto Ignition System and Lights 2 2 2 2 2 2Pitot Heater 2 - - 1 1 1 Right side may be inoperative.Alternate Static Air Source 1 1 1 1 1 1Propeller Deice System (Auto) 1 - - - - 1Propeller Deice System (Manual) 1 - - - - 1Heated Fuel Vent 2 - - - - 2Stall Warning Heater 1 - - - - 1Brake Deicer System 1 - - - - -

LANDING GEARLanding Gear Motor 1 1 1 1 1 1 May be inoperative provided operations

are continued only to a point where repairs can be accomplished.

Landing Gear Position Indicator Lights 3 3 3 3 3 3 One of three may be inoperative providedgear handle light is monitored.

Gear Handle Lights 2 2 2 2 2 2Landing Gear Aural Warning 1 1 1 1 1 1

5-13





LIGHTSCockpit and Instrument Lights * - * - * - *Lights must illuminate all instruments

and controls.Landing and Taxi Light 2 - - - - -Strobe Beacon 2 - 2 - 2 - Per FAR 91.33Position Lights 3 - 3 - 3 -Wing Ice Lights 2 - - - - * *One required for night Icing flight.Master Fault Warning Light 2 - - - - -Master Fault Caution Lights 2 - - - - -Cabin Door Caution Light 1 * * * * * *May be Inoperative provided visual Indicators

are checked prior to eachtakeoff.

NAVIGATION INSTRUMENTSAltimeter 2 1 1 1 1 1 Right side may be inoperativeAirspeed Indicator 2 1 1 1 1 1 Right side may be inoperative.Vertical Speed Indicator 2 - - - - -Standby Magnetic Compass 1 1 1 1 1 1Horizon Indicator 2 - - 2 2 2 Right side may be inoperative.Outside Air Temperature 1 1 1 1 1 1Turn and Bank Indicator 2 - - 1 1 1 Right side may be inoperative.Directional Gyro 2 - - 1 1 1 Right side may be inoperative.Clock 2 - - 1 1 1Transponder 1 - - 1 1 -Distance Measuring Equipment 1 - - - - -Navigation Equipment * - - * * * *Per AR-95-1

OXYGENOxygen System 1 1 1 1 1 1Oxygen Mask * - - - - - *Refer to oxygen requirements in Section VI.

PROPELLERSPropeller Overspeed Governor 2 2 2 2 2 2Propeller Governor Test Switch 2 2 2 2 2 2Autofeathering System 1 - - - - -Autofeathering Armed Light 2 - - - - -Reverse Not Ready Light 1 1 1 1 1 1Propeller Synchrophaser 1 - - - - -

5-14





ENGINE INDICATING INSTRUMENTSPropeller Tachometer indicator 2 2 2 2 2 2Propeller Synchroscope 1 - - - - -Gas Generator Tachometer Indicator 2 2 2 2 2 2TOT Indicator 2 2 2 2 2 2Torque Indicator 2 2 2 2 2 2

ENGINE OIL INDICATORSOil Pressure Indicator 2 2 2 2 2 2Oil Temperature Indicator 2 2 2 2 2 2Chip Detector Light 2 2 2 2 2 2

BT00226

5-15/(5-16 blank)

TM 55-1510-219-10

CHAPTER 6WEIGHT / BALANCE AND LOADING

Section I. GENERAL6-1. EXTENT OF COVERAGE.

Sufficient data has been provided so that, knowing thebasic weight and moment of the aircraft, anycombination of weight and balance can be computed.

6-2. CLASS.

Army RC-12D aircraft are in Class 1. Additionaldirectives governing weight and balance of Class 1

aircraft forms and records are contained in AR 95-3, TM55-1500-342-23, and DA PAM 738-751.

6-3. AIRCRAFT COMPARTMENT AND STATIONS.

The aircraft is separated into two compartmentsassociated with loading. These compartments are thecockpit and the cabin. Figure 6-1 shows the generaldescription of aircraft compartments.

6-1

TM 55-1510-219-10

Figure 6-1. Aircraft Compartments and Stations

6-2

TM 55-1510-219-10

Section II. WEIGHT AND BALANCE6-4. PURPOSE.

The data to be inserted on weight and balancecharts and forms are applicable only to the individualaircraft, the serial number of which appears on the titlepage of the booklet entitled WEIGHT AND BALANCEDATA supplied by the aircraft manufacturer and on thevarious forms and charts which remain with the aircraft.The charts and forms referred to in this chapter maydiffer in nomenclature and arrangement from time totime, but the principle on which they are based will notchange.

6-5. CHARTS AND FORMS.

The standard system of weight and balancecontrol requires the use of several different charts andforms. Within this chapter, the following are used:

a. Chart C Basic Weight and Balance Record, DDForm 365-3.

b. Form F Weight and Balance Clearance Form F,DD Form 365-4 (Tactical).

6-6. RESPONSIBILITY.

The aircraft manufacturer inserts all aircraftidentifying data on the title page of the booklet entitledWEIGHT AND BALANCE DATA and on the variouscharts and forms. All charts, including one sampleWeight and Balance Clearance Form F, if applicable,are completed at time of delivery. This record is thebasic weight and balance data of the aircraft at delivery.All subsequent changes in weight and balance arecompiled by the weight and balance technician.

6-7. CHART C BASIC WEIGHT AND BALANCERECORD, DD FORM 365-3.

Chart C is a continuous history of the basic weightand moment resulting from structural and equipmentchanges made in service. At all times, the last weight

and moment/100 entry is considered the current weightand balance status of the basic aircraft.

6-8. WEIGHT AND BALANCE CLEARANCE FORMF, DD FORM 365-4 (TACTICAL).

Refer to TM 55-1500-342-23 for Form Finstructions. Refer to tables 6-1, 6-2, and 6-3 formoments.

NOTEThe maximum baggagecompartment weight is 410pounds. Do not exceed the 100Lbs/Sq. Ft floor loading limit.

Table 6-1. Baggage Moment

6-3

TM 55-1510-219-10Table 6-2. Fuel Moments

6-4

TM 55-1510-219-10Table 6-3. Center of Gravity Moment Table - Moment / 100

6-5

TM 55-1510-219-10Table 6-3. Center of Gravity Moment Table - Moment / 100 (Continued)

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Table 6-3. Center of Gravity Moment Table - Moment/100 (continued)

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Section III. FUEL/OIL6-9. FUEL LOAD.

Fuel loading imposes a restriction on the amountof load which can be carried. The required fuel mustfirst be determined, then that weight subtracted from thetotal weight of passengers, baggage and fuel. Weightup to and including the remaining allowable capacity canbe subtracted directly from the weight of passengers,baggage and fuel. As the fuel load is increased, theloading capacity is reduced.

6-10. FUEL AND OIL DATA.

a. Fuel Moment Table. Table 6-2 shows fuelmoment/100 given US gallons or pounds for JP-4 andJP-5.

b. Oil Data. Total oil weight is 62 pounds and isincluded in the basic weight of the aircraft. Servicinginformation is provided in Chapter 2.

Section IV. CENTER OF GRAVITY

6-11. CENTER OF GRAVITY LIMITATIONS.

WARNINGThe forward Center of Gravity limitmay be exceeded when the missiongear is removed.

Center of gravity limitations areexpressed in ARM inches whichrefers to a positive measurementfrom the aircraft's reference datum.The forward CG limit at 11,279 Lbs.or less is 181.0 ARM inches. At14,200 Lbs. or less, the aft CG limitis 196.4 ARM inches. Tables 6-4 and6-5 provide forward and aft CGlimitations.

Table 6-4. Center of Gravity Limits (Landing Gear Down) Below 12,500 Lbs.

Section V. CARGO LOADING6-12. LOAD PLANNING.

The basic factors to be considered in any loadingsituation are as follows:

a. Cargo shall be arranged to permit access to allemergency equipment and exits during flight.

b. Floorboard structural capacity shall beconsidered in the loading of heavy or sharp-edgedcontainers and equipment. Shorings shall be used todistribute highly condensed weights evenly over thecargo areas.

Table 6-5. Center of Gravity Limits (Landing Gear Down)

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c. All cargo shall be adequately secured to preventdamage to the aircraft, other cargo, or the item itself.

6-13. LOADING PROCEDURE.

Loading of cargo is accomplished through the cabindoor (21.5 in. X 50.0 in.) or the cargo door (52.0 in. X52.0 in).

6-14. SECURING LOADS.

All cargo shall be secured with restraints strongenough to withstand the maximum force exerted in anydirection. The maximum force can be determined bymultiplying the weight of the cargo item by theapplicable load factor. These established load factors(the ratio between the total force and the weight of thecargo item) are 1.5 to the side and rear, 3.0 up, 6.6down, and 9.0 forward.

6-9/(6-10 blank)

TM 55-1510-219-10

CHAPTER 7PERFORMANCE

Section I. INTRODUCTION TO PERFORMANCE7-1. INTRODUCTION

The graphs in this section present performanceinformation for takeoff, climb, cruise, and landing atvarious parameters of weight, altitude, and temperature.

The following example presents calculations for aproposed flight from Denver to Reno using theconditions listed below:

7-2. CONDITIONS.

At Stapleton International (DEN):Free Air Temperature ..........................28°C (82°F)Field Elevation .......................................15333 feetAltimeter Setting ..................................30.25 in. HgWind ............................................. 210° at 13 knotsRunway 35R Length ........................... 112,000 feet

At Cannon International (RNO): FreeAir Temperature ..................................32°C (90°F)Field Elevation .......................................14412 feetAltimeter Setting .........................129.83 in. Hg feetWind ............................................ 260° at 10 knotsRunway 25 Length ................................16101 feet

Route segment of trip2:DEN J116 EKR J173 SLC J154BAM J32 RNO

Cruise Altitude:22,000 feet

1 Source: NOAA Standard Instrument Departures forWestern United States, 9 Jun 1983.2 Source: NOAA Enroute High Altitude -U.S.Chart H-2, 9 Jun 1983.3 MEA on NOAA Enroute Low Altitude -U.S.Chart L-8, 9 Jun 1983.4 Includes distance between airport and VORTAC, perSource in Footnote 1.

7-3. PRESSURE ALTITUDE.

To determine the approximate pressure altitude atorigin and destination airports, add 1000 feet to fieldelevation for each 1.00 in. Hg that the reportedaltimeter setting value is below 29.92 in. Hg, andsubtract 1000 feet for each 1.00 in. Hg above 29.92 in.Hg. Always subtract the reported altimeter setting from29.92 in. Hg, then multiply the answer by 1000 to findthe difference in feet between field elevation andpressure altitude.

Pressure Altitude at DEN:29.92 in. Hg 30.25 in. Hg = 0.330.33 x 1000 feet = 330 feet

The pressure altitude at DEN is 330 feet below fieldelevation.

Pressure altitude at DEN = 5333 -330 = 5003 feet

Route Segment Data 2

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Pressure altitude at RNO:29.92 in. Hg29.83 in. Hg = 0.090.09 x 1000 feet = +90 feet

The pressure altitude at RNO is 90 feet above fieldelevation.

Pressure altitude at RNO = 4412 +90 = 4502 feet.7-4. TAKEOFF WEIGHT.

Maximum takeoff weight (from limitations section)= 14,200 pounds.

7-5. TAKEOFF WEIGHT TO ACHIEVE POSITIVEONE-ENGINE-INOPERATIVE CLIMB AT LIFTOFF(FLAPS 0%).

Enter the graph at 5003 feet to 28°C, to determinethe maximum weight at which the acceleratingprocedure should be attempted.

Maximum Accelerate-Go Weight 12,990 pounds

7-6. ACCELERATE-STOP (FLAPS 0%).

Enter the accelerate-stop graph at 28°C, 5003 feetpressure altitude, 12,000 pounds, and 10 knots headwind component:

Accelerate-Stop Distance ................... 4520 feetTakeoff Decision Speed .......................99 knots

7-7 TAKE-OFF DISTANCE (FLAPS 0%).

Enter the graph at 28°C, 5003 feet pressure altitude,12,000 pounds, and 10 knots head wind component:Ground Roll ........................................ 2860 feetTotal Distance over 50-footobstacle ............................................. 4400 feetTake-off Speed:

At Rotation ...........................................99 knotsAt 50 feet ...........................................116 knots

7-8.ACCELERATE-GO FLIGHT PATH EXAMPLE.

The following example assumes the aircraft is loadedso that takeoff weight is 10,000 pounds.

a. Accelerate-Go Distance Over 50-foot Obstacle(Flaps 0%o). Enter the graph at 28°C, 5003 feetpressure altitude, 10,000 pounds, and 10 knots head windcomponent:Total Distance Over 50-foot Obstacle.......... 5710 feetSpeed at Rotation (VR) ..........................94 knotsSpeed at 35 Feet Above Runway (ClimbSpeed) ...............................................106 knots

b. Takeoff Climb Gradient One Engine Inoperative(Flaps 0%o). Enter the graph at 28°C, 5003 feetpressure altitude, and 10,000 pounds:

Climb Gradient ..........................................5.5%Climb Speed ......................................106 knots

A 5.5% climb gradient is 55 feet of vertical height per1000 feet of horizontal distance.

NOTEThe graphs for take-off climb gradientassume a zero-wind condition. Climbinginto a head wind will result in higherangles of climb, and hence betterobstacle clearance capabilities.

Calculations of the horizontal distance to clear anobstacle 100 feet above the runway surface: (fig.7-1)

Distance from 50 feet to 100 feet = 50 feet(100 50) (1000 + 55) = 910 feetTotal Distance = 5710 +910 = 6620 feet

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Figure 7-1. Accelerate-Go Flight Path

7-9. FLIGHT PLANNING.

Calculations for flight time, block speed, and fuelrequirements for a proposed flight, are detailed belowusing the same conditions presented in paragraph 7-2,and a takeoff weight of 12,000 pounds:

DENPressure Altitude ...................................... 5003 feetFAT ................................................................. 28°CISA Condition ..........................................ISA +23°C

DEN-SLCPressure Altitude ....................................22,000 feetFAT ................................................................ -20°CISA Condition ..........................................ISA +27°C

SLC-RNOPressure Altitude ....................................22,000 feetFAT ................................................................ -12°CISA Condition ........................................ISA + 1 7°C

RNOPressure Altitude ...................................... 4502 feetFAT ................................................................. 32°CISA Condition ..........................................ISA +27°C

Enter the TIME, FUEL, AND DISTANCE TOCLIMB Graph at 28°C to 5003 feet, and to 12,000pounds, and enter at -2°C to 22,000 feet, and to 12, 000pounds, and read:

Time to Climb . .......... 18.4 -2.2 = 16.2 = 16 minutesFuel Used to Climb ................ 248 -45 = 203 poundsDistance Traveled ..............50 -7 = 43 nautical miles

Enter the TIME, FUEL, AND DISTANCE TODESCEND Graph at 22,000 feet, and enter again at4502 feet, and read:Time to Descend . ....... 14.7 -3.1 =11.6 = 12 minutesFuel Used to Descend ........... 162 -39 = 123 poundsDistance Traveled . ..........67 -12 = 55 nautical miles

An estimated average cruise weight of 11,200 poundswas used for this example.

Enter the tables for MAXIMUM ENDURANCEPOWER @ 1700 RPM for ISA + 10°C, ISA +20°C, andISA + 30°C, and read the cruise speeds for 22, 000 feetat 12,000 pounds and 11,000 pounds:

Interpolate between these speeds for ISA +27°Cand ISA + 17°C at 11,200 pounds:

7-3

TM 55-1510-219-10Cruise True Airspeeds At FI 220

Cruise True Airspeed (ISA +27°C) ............ 200 knotsCruise True Airspeed (ISA +17°C) ............ 197 knots

Enter the MAXIMUM ENDURANCE POWER @1700 RPM Tables for ISA +10°C, ISA +20°C, and ISA+30°C at 12,000 pounds and 11,000 pounds andinterpolate the recommended torque settings for ISA +27°C and ISA +17°C.

ISA +27°C ............................50% torque per engineISA + 17°C ...........................49% torque per engine

Enter the MAXIMUM ENDURANCE POWER @1700 RPm, Tables for ISA + 10°C, ISA +20°C, and ISA+30°C at 12,000 pounds and 11,000 pounds at 22,000feet, and interpolate the fuel flows for ISA +27°C andISA + 17°C at 11,200 pounds.

ISA +27°CFuel Flow Per Engine ..............................237 Lbs/hrTotal Fuel Flow .......................................474 Lbs/hr

ISA +17°C

Fuel Flow Per Engine ..............................232 Lbs/hrTotal Fuel Flow .......................................464 Lbs/hr

NOTEFor flight planning, enter thesecharts at the forecasted ISAcondition; for enroute power settingsand fuel flows, enter at the actualindicated FAT.

Time and fuel used were calculated at MAXIMUMENDURANCE POWER @ 1700 RPM as follows:

Results are as follows:

7-4

TIME - FUEL - DISTANCE

TM 55-1510-219-10

7-10. RESERVE FUEL.Reserve Fuel is calculated for 45 minutes at MaximumRange Power @ 1700 RPM. Use planned cruisealtitude (22,000 feet), forecasted ISA condition (ISA +17°C), and estimated weight at end of planned trip(10,091 pounds). (Since the lowest weight column inthe tables is 11,000 pounds, assume weight at the endof the planned trip to be 11, 000 pounds, and use thatfuel flow value for this example.)

Enter the tables for MAXIMUM RANGE POWER@ 1700 RPM for ISA + 10°C and ISA +20°C at 11,000Lbs and 22,000 feet, and read the total fuel flows:

ISA +10°C ...............................................468 Lbs/hrISA +20°C................................................484 Lbs/hr

Then interpolate for the fuel flow at ISA + 17°C asfollows: Change in Fuel Flow = 484 468 = 16Lbs/hr.Change in Temperature = (ISA +20°C) (ISA + 10C) = 10°C.Rate of Change in Fuel Flow = Change in FuelFlow + Change in Temperature.Rate of Change in Fuel Flow = (16 Lbs/hr) +(10°C).Rate of Change in Fuel Flow = 1.6 Lbs/hr increaseper 1°C increase.Temperature increase from ISA + 10°C to ISA+17°C = 7°C.Total Change in Fuel Flow = 7 X 1.6 Lbs/hr = 11.2Lbs/hr.Total Fuel Flow = (ISA +10°C Fuel Flow) + (TotalChange in Fuel Flow).Total Fuel Flow = (468) + (11.2) = 479.2 Lbs/ hr.Reserve Fuel = 45 minutes X Total Fuel Flow.Reserve Fuel = (0.75) X (479.2 Lbs/hr) = 359.4 =360 lbs/hr.Total Fuel Requirement = 1999 +360 = 2359pounds.

7-11. ZERO FUEL WEIGHT LIMITATION.For this example, the following conditions were

assumed:Ramp Weight ................................... 12,090 poundsWeight of Usable Fuel Onboard ..........2359 Pounds

Zero Fuel Weight = Ramp Weight - Weight ofUsable Fuel Onboard

Zero Fuel Weight = (12,090) (2359) = 9731poundsMaximum zero fuel weight limitation (fromChapter 5) = 11,500 pounds.Maximum Zero Fuel Weight Limitation has notbeen exceeded.

Anytime the Zero Fuel Weight exceeds theMaximum Zero Fuel Weight Limit, the excess must beoff-loaded from PAYLOAD. If desired, additional FUELONLY may then be added until the ramp weight equalsthe Maximum Ramp Weight Limit of 14,290 Lbs.

7-12. LANDING INFORMATION.

The estimated Landing Weight is determined bysubtracting the fuel required for the trip from the RampWeight:Ramp Weight..........................................12,090 LbsFuel Required for Total Trip ................ 1999 poundsLanding Weight (12,090 - 1999) . ...... 10,091 pounds

Enter the NORMAL LANDING DISTANCEWITHOUT PROPELLER REVERSING FLAPS 100%Graph at 32°C, 4502 feet, 10,091 pounds, and 10 knotshead wind component:Ground Roll............................................... 1725 feetTotal Distance Over 50-foot Roll .............. 1725 feetTotal Distance Over 50-foot Obstacle ....... 3060 feetApproach Speed ...................................... 100 knots

Enter the CLIMB - BALKED LANDING Graph at32°C, 4502 feet, and 10,091 pounds:

Rate of Climb .........................................1550 ft/minClimb Gradient .............................................. 12.5%7-13. COMMENTS PERTINENT TO THE USE OFPERFORMANCE GRAPHS.

a. In addition to presenting the answer for aparticular set of conditions, the example on a graph alsopresents the order in which the various scales on thegraph should be used. For instance, if the first item inthe example is FAT, then enter the graph at the existingFAT.

b. The reference lines indicate where to beginfollowing the guidelines. Always project to the referenceline first, then follow the guidelines to the next knownitem by maintaining the same PROPORTIONAL

7-5

TM 55-1510-219-10

DISTANCE between the guide line above and the guideline below the projected line. For instance, if theprojected line intersects the reference line in the ratio of30% down/70% up between the guidelines, thenmaintain this same 30%/70% relationship between theguide lines and follow them to the next known item.

c. The associated conditions define thespecific conditions from which performance parametershave been determined. They are not intended to beused as instructions; however, performance valuesdetermined from charts can only be achieved if thespecified conditions exist.

d. The full amount of usable fuel is availablefor all approved flight conditions.

e. Notes have been provided on variousgraphs and tables to approximate performance with icevanes extended. The effect will vary, depending uponairspeed, temperature, altitude, and ambient conditions.At lower altitudes, where operation on the torque limit ispossible, the effect of ice vane extension will be less,depending upon how much power can be recoveredafter the ice vanes have been extended.

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TM 55-1510-219-10

Index of FiguresAccelerate Go Flight Path ................................................................................................................................7-3Airspeed Calibration - Normal System ............................................................................................................7-10Airspeed Calibration - Alternate System .........................................................................................................7-11Altimeter Correction - Normal System ............................................................................................................7-12Altimeter Correction - Alternate System .........................................................................................................7-13Indicated Outside Air Temperature Correction ................................................................................................7-14ISA Conversion ..............................................................................................................................................7-15Fahrenheit to Celsius Temperature Conversion ..............................................................................................7-16Take-Off Weight - Flaps 0% - To Achieve Positive One-Engine Inoperative Climb at Lift-Off .........................7-17Take-Off Weight - Flaps 40% - To Achieve Positive One-Engine Inoperative Climb at Lift-Off .......................7-18Wind Components .........................................................................................................................................7-19Take-Off Distance - Flaps 0% ........................................................................................................................7-20Take-Off Distance - Flaps 40% ......................................................................................................................7-21Minimum Take-Off Power at 2000 RPM .........................................................................................................7-22Accelerate-Stop - Flaps 0%.............................................................................................................................7-23Accelerate-Stop - Flaps 40% ..........................................................................................................................7-24Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 0% ...............................................................................7-25Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 40% .............................................................................7-26Time, Fuel, and Distance to Cruise Climb ......................................................................................................7-27Climb Two-Engines Flaps 0%..........................................................................................................................7-28Climb Two-Engines Flaps 40%........................................................................................................................7-29Take-Off Climb Gradient - One-Engine Inoperative - Flaps 0% ......................................................................7-30Take-Off Climb Gradient - One-Engine Inoperative - Flaps 40% ....................................................................7-31Climb One-Engine Inoperative .......................................................................................................................7-32Climb - Balked Landing ...................................................................................................................................7-33Service Ceiling One-Engine Inoperative..........................................................................................................7-34Maximum Cruise Power @ 1900 RPM, ISA -30 ....................................................................................7-35, 7-36Maximum Cruise Power @ 1900 RPM, ISA -20 . ...................................................................................7-37, 7-38Maximum Cruise Power @ 1900 RPM, ISA -10° ...................................................................................7-39, 7-40Maximum Cruise Power @ 1900 RPM, ISA ..........................................................................................7-41, 7-42Maximum Cruise Power @ 1900 RPM, ISA +10° ..................................................................................7-43, 7-44Maximum Cruise Power @ 1900 RPM, ISA +20° . .................................................................................7-45, 7-46Maximum Cruise Power @ 1900 RPM, ISA +30° ..................................................................................7-47, 7-48Maximum Cruise Power @ 1900 RPM, ISA +37° ..................................................................................7-49, 7-50Maximum Cruise Power @ 1900 RPM ...........................................................................................................7-51Maximum Cruise Speeds @ 1900 RPM .........................................................................................................7-52Fuel Flow At Maximum Cruise Power @ 1900 RPM, ......................................................................................7-53

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Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA -30°C ............................................7-54,7-55Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA -20°C ............................................7-56,7-57Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA -10°C ............................................7-58,7-59Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA ......................................................7-60,7-61Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +10°C ...........................................7-62,7-63Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +20°C ...........................................7-64,7-65Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +30°C ...........................................7-66,7-67Maximum Cruise Power One-Engine Inoperative @ 1900 RPM, ISA +37°C ...........................................7-68,7-69Maximum Cruise Power @ 1700 RPM, ISA -30°C ..................................................................................7-70,7-71Maximum Cruise Power @ 1700 RPM, ISA -20°C ..................................................................................7-72,7-73Maximum Cruise Power @ 1700 RPM, ISA -10°C ..................................................................................7-74,7-75Maximum Cruise Power @ 1700 RPM, ISA ...........................................................................................7-76.7-77Maximum Cruise Power @ 1700 RPM, ISA +10°C .................................................................................7-78,7-79Maximum Cruise Power @ 1700 RPM, ISA +20°C .................................................................................7-80,7-81Maximum Cruise Power @ 1700 RPM, ISA +30°C .................................................................................7-82,7-83Maximum Cruise Power @ 1700 RPM, ISA +37°C .................................................................................7-84,7-85Maximum Cruise Power @ 1700 RPM ...........................................................................................................7-86Maximum Range Power @ 1700 RPM, ISA -30°C .................................................................................7-87,7-88Maximum Range Power @ 1700 RPM, ISA -20°C .................................................................................7-89,7-90Maximum Range Power @ 1700 RPM, ISA -10°C .................................................................................7-91,7-92Maximum Range Power @ 1700 RPM, ISA ............................................................................................7-93,7-94Maximum Range Power @ 1700 RPM, ISA +10°C ................................................................................7-95,7-96Maximum Range Power @ 1700 RPM, ISA +20°C ................................................................................7-97,7-98Maximum Range Power @ 1700 RPM, ISA +30°C .............................................................................. 7-99,7-100Maximum Range Power @ 1700 RPM, ISA +37°C ............................................................................ 7-101,7-102Maximum Endurance Power @ 1700 RPM, ISA -30°C ....................................................................... 7-103,7-104Maximum Endurance Power @ 1700 RPM, ISA -20°C ....................................................................... 7-105,7-106Maximum Endurance Power @ 1700 RPM, ISA -10°C ....................................................................... 7-107,7-108Maximum Endurance Power @ 1700 RPM, ISA ................................................................................. 7-109,7-110Maximum Endurance Power @ 1700 RPM, ISA + 10°C ..................................................................... 7-111,7-112Maximum Endurance Power @ 1700 RPM, ISA + 20°C ..................................................................... 7-113,7-114Maximum Endurance Power @ 1700 RPM, ISA +30°C ...................................................................... 7-115,7-116Maximum Endurance Power @ 1700 RPM, ISA +37°C ...................................................................... 7-117,7-118Range Profile - Maximum Cruise Power @ 1900 RPM .................................................................................7-119Range Profile - Maximum Range Power @ 1700 RPM .................................................................................7-120Range Profile - 542 Gallons Usable Fuel.......................................................................................................7-121Endurance Profile - 542 Gallons Usable Fuel ...............................................................................................7-122Time, Fuel, And Distance to Descend ..........................................................................................................7-123

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Landing Distance Without Propeller Reversing - Flaps 0% ...........................................................................7-124Landing Distance Without Propeller Reversing - Flaps 100% ......................................................................7 -125Landing Distance With Propeller Reversing - Flaps 0% ................................................................................7-126Landing Distance With Propeller Reversing - Flaps 100% ............................................................................7-127Stopping Distance Factors ...........................................................................................................................7-128

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Figure 7-2. Airspeed Calibration - Normal System

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Figure 7-3. Airspeed Calibration - Alternate System

7-11

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Figure 7-4. Altimeter Correction - Normal System

7-12

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Figure 7-5. Altimeter Correction - Alternate System7-13

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Figure 7-6. Indicated Outside Air temperature Correction

7-14

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Figure 7-7. ISA Conversion

7-15

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Figure 7-8. Fahrenheit to Celsius Temperature Conversion

7-16

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Figure 7-9. Take-Off Weight - Flaps 0%, To Achieve Positive One-Engine Inoperative Climb at Lift-Off

7-17

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Figure 7-10. Take-Off Weight - Flaps 40%, To Achieve Positive One-Engine Inoperative Climb at Lift-Off

7-18

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Figure 7-11. Wind Components

7-19

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Figure 7-12. Takeoff Distance - Flaps 0%

7-20

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Figure 7-13. Take-Off Distance - Flaps 40%

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MINIMUM TAKEOFF POWER AT 2000 RPM(65 KNOTS)

NOTES: 1. TORQUE INCREASES APPROXIMATELY 1% FROM 0 TO 65 KNOTS.

2. THE PERCENT TORQUE INDICATED IN THIS FIGURE IS THE MINIMUM VALUE AT WHICHTAKE-OFF PERFORMANCE PRESENTED IN THE SECTION CAN BE REALIZED. ANY EXCESSPOWER WHICH MAY BE DEVELOPED WITHOUT EXCEEDING ENGINE LIMITATIONS MAY BEUTILIZED.

Figure 7-14. Minimum Take-Off Power at 2000, RPM

7-22

Figure 7-15. Accelerate-Stop - Flaps 0%7-23

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Figure 7-16. Accelerate-Stop - Flaps 40

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Figure 7-17. Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 0%

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Figure 7-18. Accelerate-Go Distance Over 50-Ft Obstacle - Flaps 40%

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Figure 7-19. Time, Fuel, and Distance to Cruise Climb

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Figure 7-20. Climb - Two Engines - Flaps 0%

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Figure 7-21. Climb - Two Engines - Flaps 40%

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Figure 7-22. Take-Off Climb Gradient - One-Engine Inoperative - Flaps 0%

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Figure 7-23. Take-Off Climb Gradient - One-Engine Inoperative - Flaps 40%

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Figure 7-24. Climb - One Engine Inoperative

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Figure 7-25. Climb - Balked Landing

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Figure 7-26. Service Ceiling - One Engine Inoperative

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Figure 7-27. Maximum Cruise Power 1900 RPM, ISA -30°C (Sheet 1 of 2)

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Figure 7-29. Maximum Cruise Power 1900 RPM, ISA - 10°C (Sheet 1 of 2)

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Figure 7-30. Maximum Cruise Power 1900 RPM, ISA (Sheet 1 of 2)

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Figure 7-31. Maximum Cruise Power 1900 RPM, ISA + 10°C (Sheet 1 of 2)

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Figure 7-32. Maximum Cruise Power 1900 RPM, ISA +20°C (Sheet 1 of 2)

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7-46

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7-47

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7-48

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7-49

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Figure 7-35. Maximum Cruise Speeds @ 1900 RPM

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Figure 7-36. Maximum Cruise Power @ 1900 RPM

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Figure 7-37. Fuel Flow at Maximum Cruise Power @ 1900 RPM

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Figure 7-38. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -30°C (Sheet 1 of 2)

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Figure 7-38. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -30°C (Sheet 2 of 2)7-55

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Figure 7-40. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA - -10°C (Sheet 1 of 2)

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Figure 7-40. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA -10°C (Sheet 2 of 2)

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Figure 7-41. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA (Sheet 1 of 2)

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Figure 7-41. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA (Sheet 2 of 2)

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Figure 7-42. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA + 10°C (Sheet 1 of 2)

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Figure 7-42. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA + 10° C (Sheet 2 of 2)

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Figure 7-43. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +20°C (Sheet 1 of 2)

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Figure 7-44. Maximum Cruise Power One-Engine Inoperative 1900 RPM, ISA +30°C (Sheet 1 of 2)7-66

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7-73

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7-74

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7-75

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7-76

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7-77

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Figure 7-50. Maximum Cruise Power 1700 RPM, ISA ± 10°C (Sheet 1 of 2)

7-78

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7-79

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7-80

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7-81

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7-82

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MAXIMUM CRUISE POWER

1700 RPM

WEIGHT: 12,000 LBS

NOTES: 1. ISA DEVIATION LINES REFLECT ACTUAL TEMPERATURES FOR FLIGHT PLANNING. INDICATEDTEMPERATURES SHOULD BE USED FOR IN-FLIGHT CRUISE POWER SETTINGS.

2. FOR OPERATION WITH ICE VANES EXTENDED, TORQUE WILL DECREASE APPROXIMATELY20%. IF DESIRED, ORIGINAL POWER MAY BE RESET, PROVIDED ITT LIMIT IS NOT EXCEEDED.

Figure 7-54. Maximum Cruise Power @ 1700 RPM

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Figure 7-55. Maximum Range Power @ 1700 RPM, ISA-30°C (Sheet 1 of 2)

7-87

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7-88

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Figure 7-56. Maximum Range Power @ 1700 RPM, ISA-200C (Sheet 1 of 2)

7-89

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7-90

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7-91

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Figure 7-57. Maximum Range Power @ 1700 RPM, ISA- 10°C (Sheet 2 of 2)

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Figure 7-58. Maximum Range Power @ 1700 RPM, ISA (Sheet 1 of 2)

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Figure 7-58. Maximum Range Power @ 1700 RPM, ISA (Sheet 2 of 2)

7-94

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Figure 7-59. Maximum Range Power @ 1700 RPM, ISA + 10°C (Sheet 1 of 2)

7-95

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Figure 7-59. Maximum Range Power @ 1700 RPM, ISA + 10°C (Sheet 2 of 2)

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Figure 7-60. Maximum Range Power @ 1700 RPM, ISA +20°C (Sheet 1 of 2)

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Figure 7-61. Maximum Range Power @ 1700 RPM, ISA+30°C (Sheet 1 of 2)

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Figure 7-61. Maximum Range Power @ 1700 RPM, ISA +30° C (Sheet 2 of 2)

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7-101

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7-102

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Figure 7-63. Maximum Endurance Power @ 1700 RPM, ISA-30°C (Sheet 1 of 2)

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7-104

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7-105

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7-106

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7-107

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Figure 7-65. Maximum Endurance Power @ 1700 RPM, ISA- 10°C (Sheet 2 of 2)

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Figure 7-66. Maximum Endurance Power @ 1700 RPM, ISA (Sheet 1 of 2)

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Figure 7-66. Maximum Endurance Power @ 1700 RPM, ISA (Sheet 2 of 2)

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Figure 7-67. Maximum Endurance Power @ 1700 RPM, ISA + 10°C (Sheet 1 of 2)

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Figure 7-67. Maximum Endurance Power @ 1700 RPM, ISA + 10°C (Sheet 2 of 2)

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Figure 7-68. Maximum Endurance Power @ 1700 RPM, ISA +20°C (Sheet 1 of 2)

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Figure 7-68. Maximum Endurance Power @ 1700 RPM, ISA +20 °C (Sheet 2 of 2)

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7-116

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Figure 7-70. Maximum Endurance Power @ 1700 RPM, ISA +37°C (Sheet I of 2)

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Figure 7-71. Range Profile-Maximum Cruise Power @ 1900 RPM7-119

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Figure 7-72. Range Profile-Maximum Range Power @ 1700 RPM

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Figure 7-73. Range Profile-542 Gallons Usable Fuel7-121

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Figure 7-74. Endurance Proflie-542 Gallons Usable Fuel7-122

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Figure 7-75. Time, Fuel, and Distance to Descend

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Figure 7-76. Landing Distance Without Propeller Reversing-Flaps 0%

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Figure 7-77. Landing Distance Without Propeller Reversing-Flaps 100%7-125

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Figure 7-78. Landing Distance With Propeller Reversing-Flaps 0%

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Figure 7-79. Landing Distance With Propeller Reversing-Flaps 100%

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Figure 7-80. Stopping Distance Factors

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CHAPTER 8NORMAL PROCEDURES

Section I. MISSION PLANNING

8-1. MISSION PLANNING.Mission planning begins when the mission is

assigned and extends to the preflight check of theaircraft. It includes, but is not limited to, checks ofoperating limits and restrictions; weight, balance, andloading; performance; publications; flight plan; and crewand passenger briefings. The pilot in command shallinsure compliance with the contents of this manual thatare applicable to the mission.

8-2. CREW BRIEFINGS.A crew briefing must be conducted for a thorough

understanding of individual and team responsibilities.The briefing should include, but not be limited to,copilot, crew chief, and ground crew responsibilities andthe coordination necessary to complete the missionmost efficiently. A review of visual signals is desirablewhen ground guides do not have a direct voicecommunications link with the crew. Refer to Section VIfor crew briefings.

Section II. OPERATING PROCEDURES AND MANEUVERS

8-3. OPERATING PROCEDURES AND MANEUVERS.This section deals with normal procedures and

includes all steps necessary for safe and efficientoperation of the aircraft from the time, a preflight beginsuntil the flight is completed and the aircraft is parkedand secured. Unique feel, characteristics, and reactionof the aircraft during various phases of operation andthe techniques and procedures used for taxiing, takeoff,climb, etc., are described, including precautions to beobserved. Only the duties of the minimum crewnecessary for the actual operation of the aircraft areincluded.

8-4. CHECKLIST.Normal procedures are given primarily in checklist

form and are amplified as necessary in accompanyingparagraph form when a detailed description of aprocedure or maneuver is required. A condensedversion of the amplified checklist, omitting allexplanatory text, is contained in the Operator's andCrewmember's Checklist, TM 55-1510-219-CL. Toprovide for easier cross referencing, the proceduralsteps are numbered to coincide with the correspondingnumbered steps in TM 55-1510-219-CL.

8-5. SYMBOLS DEFINITION.Items which apply only to night or only to

instrument flying shall have an "N" or "I", respectively,immediately preceding the check to which it is pertinent.

The symbol "O" shall be used to indicate "if installed."The star symbol indicates an operational checkcontained in the performance section of the condensedchecklist. The asterisk symbol "*" indicates thatperformance of the step is mandatory for all thru-flights.The asterisk applies only to checks performed prior totakeoff. Placarded items appear in upper case.

8-6. BEFORE EXTERIOR CHECK.1. Publications-Check DA Forms 2408-12, -13, -

14, and -18, DD Form 365-4, locally requiredforms and publications, and availability ofoperator's manual (-10) and checklist (-CL).

2. Oxygen system-Check as required.

a. Oxygen supply pressure gages-Check.

b. Supply control lever (green)-ON.

c. Diluter control lever-100% OXYGEN.

d. Emergency control lever (red)-Set toTEST MASK position while holding maskdirectly away from face, then return toNORMAL.

e. Oxygen mask-Don and adjust.

f. Emergency control lever (red)-Set toTEST MASK position and check mask forleaks, then return lever to NORMAL.

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Figure 8-1. Exterior Inspection

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g. Flow indicator-Check, during inhalationblinker appears, during exhalation blinkerdisappears). Repeat a minimum of 3times.

3. Flight controls-Unlock and check.4. Parking brake-Set.5. Manual trim-Zero.6. Gear-Down.7. Ice vanes-IN.8. Overhead panels witches and circuit breakers-

Set as follows.a. Circuit breakers-Check.b. Light dimming controls-As required.c. Cabin temperature mode-OFF.d. Ice and rain switches-As required.e. Exterior light switches-As required.f. Master panel light switch-As required.g. Inverter switches-OFF.h. Avionics master power switch-As

required.i. Environmental switches-As required.j. Autofeather switches-OFF.k. #1 ignition and engine start switch-OFF.l. Master switch-OFF.

m. #2 ignition and engine start switch-OFF.n. Standby pump switches-OFF.o. Auxiliary transfer switches-AUTO.p. Crossfeed switch-OFF.

9. AC and DC GPU's-As required.10. External power advisory lights-As required.11. Keylock switch-ON.12. Fuel pumps/crossfeed operation-Check as

follows:a. Fire pull handles-Pull.b. Standby pump switches-ON.c. Battery switch-ON.d. #1 and#2 fuel press warning lights-

Illuminated.e. Fire pull handles-IN.

f. #1 and#2 fuel pressure warning lights-Extinguished.

g. Standby pump switches-STANDBYPUMP.

h. #1 and #2 fuel pressure warning lights-Illuminated.

i. Crossfeed-Check. Check systemoperation by activating switchmomentarily left then right, noting that#1/#2 FUEL PRESS warning lightsextinguish and that the FUELCROSSFEED advisory light illuminatesas switch is energized.

13. DC power-Check (24 VDC minimum for battery,28 maximum for GPU starts).

14. Lighting systems-Check.15. Anti-ice systems-Check.

a. Stall warning heat switch-ON.b. Pitot heat switches (2)-ON. Check cover

removed.c. Fuel vent heat switches (2)-ON.d. Left wing heated fuel vent-Check by feel

for heat and condition.e. Stall warning vane-Check by feel for heat

and condition.f. Left pitot tube-Check by feel for heat and

free of obstructions.g. Right pitot tube-Check by feel for heat

and free of obstructions.h. Right wing heated fuel vent-Check by feel

for heat and condition.i. Stall warning heat switch-OFF.j. Pitot heat switches (2)-OFF.k. Heated fuel vent switches (2)-OFF.

16. Annunciator panels-Test as required.a. MASTER CAUTION, MASTER

WARNING, #1 FUEL PRESSURE, #2FUEL PRESSURE, GEAR DN, L BL AIRFAIL, R BL AIR FAIL, ALT WARN, INSTAC, #1 DC GEN, #1 INVERTER, #1 NOFUEL XFR, #2 NO FUEL XFR, #2INVERTER, #2 DC GEN-Check on.

b. ANNUNCIATOR TEST switch-Press andhold. Check that all lights in aircraft andmission annunciator panels illuminate,FIRE PULL handle lights, marker beaconlights, MASTER CAUTION and MASTERWARNING

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TM 55-1510-219-10

lights are on. Release switch and checkthat all lights except those in step (a.) areextinguished.

c. MASTER CAUTION and MASTERWARNING lights-Press. Both lightsextinguish.

d. Stall and gear warning system-TEST.Check that warning horn sounds and thatthe LDG gear control handle lights (2)illuminate.

17. Fire protection system-Check as follows:a. Fire detector test switch-Rotate

counterclockwise to check three DETRpositions. FIRE PULL handles shouldilluminate in each position. ResetMASTER WARNING in each position.

b. Fire detector test switch-Rotatecounterclockwise to check two EXTGHpositions. SQUIB OK light, associatedEXTGH DISCH caution light andMASTER CAUTION LIGHT shouldilluminate in each position.

18. INS alignment-As required.a. Exterior power 28 VDC-Connected.b. Key lock switch-ON.c. Battery switch-ON.d. Aircraft inverters-ON.e. Mission inverters-ON.f. Mission control switch-As required.g. 3 phase A.C. bus-RESET.

Check inertial cooling for ON.h. Aircraft master avionics switch-

EXTERNAL POWERi. Mode selector-ALIGNj. Data selector-DSTRK/STS

Align condition is shown on 5 digit RH display.Align will not progress beyond 8 until present position isloaded.

k. Test button-PRESS AND HOLD.All displays read 8, ROLL LIMIT, HOLD,

INSERT/ADVANCE, WY PT, CHG, ALERT, BAT,WARN, and READY NAV LAMPS lit. When released,all extinguish except INSERT/ADVANCE. Insure allmalfunction codes are cleared.

l. Data selector-L/L POS (UTM for gridnav).

m. Load present position.(1) Select N or S degrees, minutes and

tenths -INSERT/ADVANCE -PRESS.

(2) Select E or W degrees, minutes andtenths -INSERT/ADVANCE-PRESS.

NOTEInsure correct values for UTM gridspheroid coefficients are loadedusing UTM coordinates.

WAYPOINT SELECTIONa. Data selector-L/L WY PT.b. WYPT thumb wheel-DESIRED WY PT

(Do not use 0).c. LAT/LONG waypoint (Deg., Min., Tenths)-

LOAD.d. INSERT ADVANCE-PRESS AGAIN.

Latitude and longitude in arc-seconds relating totenths entered is shown.

e. ARC/SEC LAT & LONG (Sec. andTenths)-LOAD.

Repeat a thru e for each WY PT.f. Flight plan cross check-Data selector to

DIS/TIME (left display will indicatedistance between WY PTS TO-FROM).Press WY PT change. Verify logicaldistance between waypoints.

TACAN STATION SELECTIONa. Data selector-L/L WY PT.b. Simultaneously press-Keys 7 and 9.c. WY PT thumb wheel-DESIRED TACAN.d. TACAN station position-LOAD

LAT/LONG.e. INSERT/ADVANCE-PRESS.

Latitude and longitude in arc-seconds relating to tenthsentered is shown.

f. ARC/SEC and Tenths-LOAD.g. INSERT/ADVANCE-PRESS.h. TACAN station altitude-LOAD (select Key

4 or 6). Enter altitude.i. INSERT/ADVANCE-PRESS.

8-4

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j. TACAN channel number-LOAD (selectKey 2 or 8). Enter channel.

Repeat steps a thru j for all TACAN stations.HSI INTERFACE TEST

NOTEInterface test must be performedafter alignment progresses fromstate 8 but prior to switching toNAV.a. AUTO/MAN switch-MAN.

b. Couple INS to FD and engage AP.

c. Data selector (any position but DSRTK).

d. CDU TEST-PRESS and HOLD.MSU and CDU all lamps lit 8's. RMI all angles are

30 degrees. Cross track deviation bar is 1 dot right,NAV flags are retracted, WX radar NAV displayindicates 1 dot right of course.

On HRI, a 15 degree RH steering command.Aircraft panel LINK UPDATE/TACAN UPDATEannunciated and INS light illuminated.

e. Continue holding TEST switch from MANto AUTO. RMI all angles are 0. Crosstrack deviation is one dot left. A 15degree left steering command is issued.

f. Release TEST switch-Operations return tonormal.

19. Electric elevator trim and autopilot/flight directoroperation-Check as follows:

a. Pilot and copilot PITCH TRIM switches-Press to NOSE UP and NOSE DNpositions, singularly and in pairs. Checkthat trim wheel moves in proper directionand operates only when trim switches arepressed in pairs. Any deviation requiresthat electric elevator trim be turned offand flight conducted using manual trim.

b. TRIM DISC switch-Press and check thatelectric trim disconnects and that ELEVTRIM light extinguishes.

c. Flight director (FD) and radio magneticindicator (RMI) warning flags masked-Check.

NOTESince the pressure of airflow thatnormally opposes movement ofcontrol surfaces is absent duringpreflight check, it is possible toget a hard over control surface

deflection if an autopilotcommand is allowed to remainactive for any appreciable lengthof time. Move turn knob and pitchthumbwheel only enough tocheck operation, then return themto the center position.d. Select HDG mode-Check.

e. Horizontal situation indicator (HSI)heading marker under lubber line-Set.

f. Engage autopilot and check controls stiff,and AIL HI TORQUE, HDG, and AP ENGare illuminated-Check.

g. Move HSI heading marker 10° left andright and verify that FD and controlwheels respond in the appropriatedirection-Check.

h. Press AP/YD disengage switch and verifythat autopilot disengages and that flightcontrols are free -Check.

i. Engage autopilot-Check.

j. Command 5° trim UP with AP pitch wheeland verify that manual trim wheel movesnose UP and AP trim light indicates UPtrim-Check.

k. Press pitch trim switch nose down andverify that autopilot disengages andAUTOPILOT TRIM FAIL and MASTERWARNING lights illuminate-Check.

NOTEThe AP TRIM FAIL annunciatorwill extinguish by pressing theAP/YD disconnect button on thecontrol wheel to the first detent.l. Repeat steps i thru k above using

opposite commands.

m. Engage autopilot-Check.

n. Move HSI heading marker to command abank on flight director-Check.

o. Press GO-AROUND switch and verify thatGA annunciator light illuminates, autopilotdisengages, and that flight directorcommands a wings level, 7° nose-upattitude-Check.

p. Press TEST switch (pilot's HRI) and verifythat attitude display indicates an

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TM 55-1510-219-10

additional 10° pitch up and 20° right bank-Check.

q. Engage autopilot command DN with APpitch wheel and engage and hold AUTOPILOT TRIM TEST switch when elevatortrim wheel starts to rotate.

r. Verify that autopilot disengages and APTRIM FAIL and MASTER WARNINGlights illuminate within 10 seconds.

20. Avionics-Check as follows:a. NAV 1

(1) Frequency select knob (NAV panel)-Select a VOR frequency.

(2) NAV TEST switch (NAV PANEL)-Press and hold.

(3) RMI-Observe that single needleindicates approximately 005°.

(4) VOR/LOC flag -Check that flag isout of view.

(5) TO/FROM pointer-Check thatpointer indicates TO.

(6) HSI course deviation bar-Check forcentered bar.

(7) Marker beacon lights-Check that allthree lamps are illuminated andflickering at approximately a 30 Hzrate.

(8) VOR frequency knob (NAV panel)-Select a LOC frequency.

(9) HSI course deviation bar-Check thatbar indicates a deflection ofapproximately one dot right ofcenter.

(10) HSI glideslope pointer-Check thatpointer indicates a deflection ofapproximately one dot below center.

(11) Marker beacon lights-Check that allthree lamps are illuminated andflickering at approximately a 30 Hzrate.

b. NAV 2

All NAV 2 self-test procedures are the same asthose used for NAV 1, with the exception of the markerbeacon test. There is no marker beacon receiver in theNAV 2 system.

c. TACAN(1) TEST pushbutton -Press and hold.(2) Range indicator-Check for an

indication of 0.0 ± 0.1 nauticalmiles.

(3) Pilot's COURSE SELECTORswitch-Select TACAN.

(4) Pilot's RMI selector switch-SelectTACAN.

(5) RMI double needle-Check for anindication of 180° ± 2°.

(6) HSI course selector-Turn to 180°and adjust slowly until the coursedeviation bar is centered. The barshould center between a selectedcourse of 178° to 182°.

(7) HSI course selector-Turn theselector + 10° from the settingachieved in step 6, and check thatcourse deviation bar is located overthe far left 10° dot.

(8) HSI course selector-Turn theselector + 10° from the setting

(9) HSI course selector-Turn theselector -10° from the settingachieved in step 6, and check thatcourse deviation bar is located overthe far right 10° dot.

(10) TO-FROM indicator-Check that TOis indicated.

(11) TEST pushbutton-Release.21. Flaps-Check.22. Battery switch-As required.23. Toilet-Check.24. Emergency equipment-Check.25. Mission equipment and circuit breakers-Check

and set.26. Parachutes-Check (as required).

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TM 55-1510-219-10

8-7. FUEL SAMPLE.NOTE

Fuel and oil quantity check maybe performed prior to BEFOREEXTERIOR CHECK. During warmweather open fuel cap slowly toprevent being sprayed by fuelunder pressure due to thermalexpansion.

*1. Check collective fuel sample from all drains forpossible contamination. Thru-flight check isonly required if aircraft has been refueled.

8-8. LEFT WING, AREA 1.a. Left wing area-Check as follows (fig. 8-1):

*1. General condition-Check for skin damagesuch as buckling, splitting, distortion,dents, or fuel leaks.

2. Flaps-Check for full retraction(approximately 0.25 inch play) and skindamage such as buckling, splitting,distortion, or dents.

3. Fuel sump drains (3)-Check for leaks.4. Controls and trim tab-Check security and

trim tab rig.NOTE

All static wicks (24) must beinstalled for optimum radioperformance.5. Static wicks-Check security and condition.6. Wing pod, navigation lights and antennas

(2) -Check condition.7. Recognition light-Check condition.8. Outboard antenna set-Check condition.

*9. Main tank fuel and cap-Check fuel levelvisually, condition of seal, and cap tightand properly installed.

10. Outboard wing fuel vent-Check free ofobstructions.

11. Outboard deice boot-Check for securebonding, cracks, loose patches, stallstrips, and general condition.

12. Stall warning vane-Check free.13. Tiedown-Released.14. Inboard dipole antenna set-Check for

security and cracks at mounting points.Check bonding secure, boots free of cutsand cracks.

15. Wing ice light-Check condition.16. AC GPU access door-Secure.17. Recessed and heated fuel vents-Check

free of obstruction.18. Inverter inlet and exhaust louvers-Check

free of obstructions.8-9. LEFT MAIN LANDING GEAR.

a. Left main landing gear-Check as follows:1. Tires-Check for cuts, bruises, wear,

proper inflation and wheel condition.2. Brake assembly-Check brake lines for

damage or signs of leakage, brake liningsfor wear (0.25-inch), brake deiceassembly and bleed air hose for conditionand security.

3. Shock strut-Check for signs of leakage,minimum strut extension, (5.50 inches)and that left and right strut extension isapproximately equal.

4. Torque knee-Check condition.5. Safety switch-Check condition, wire, and

security.6. Fire extinguisher pressure-Check

pressure within limits.7. Wheel well, doors, and linkage-Check for

signs of leaks, broken wires, security, andgeneral condition.

8. Fuel sump drains (forward)-Check forleaks.

8-10. LEFT ENGINE AND PROPELLER.a. Left engine-Check as follows:

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TM 55-1510-219-10

CAUTIONA cold oil check is unreliable. Oilshould be checked within 10minutes after stopping engine. Ifmore than 10 minutes haveelapsed, motor engine for 30seconds, then recheck. If morethan 10 hours have elapsed, runengine for 2 minutes, thenrecheck. Add oil as required. Donot overfill.*1. Engine oil-Check oil level, oil cap secure,

locking tab aft, and access door locked.NOTE

Secure front cowling latches first.2. Engine compartment, left side-Check for

fuel and oil leaks, security of oil cap, doorlocking pins, and general condition.

*3. Left cowl locks-Locked.4. Left exhaust stack-Check for cracks,

security and free of obstructions.*5. Propeller blades and spinner-Check blade

condition, boots, security of spinner andfree propeller rotation.

*6. Engine air inlets and ice vane-Check freeof obstruction and ice vane retracted.

7. Bypass door-Check condition.*8. Right cowl locks-Locked.9. Right exhaust stack-Check for cracks,

security and free of obstructions.10. Engine compartment, right side-Check

for fuel and oil leaks, ice vane linkage,door locking pins, and general condition.Lock compartment access door.

8-11. CENTER SECTION, LEFT SIDE.a. Center section-Check as follows:

1. Heat exchanger inlet and outlet-Check forcracks and free of obstruction.

2. Auxiliary tank fuel sump drain-Check forleaks.

3. Deice boot-Check for secure bonding,cracks, loose patches, and generalcondition.

*4. Auxiliary tank fuel gage and cap-Checkfuel level visually, condition of seal, andcap tight and properly installed.

5. Monopole antenna-Check condition.8-12. FUSELAGE UNDERSIDE.

a. Fuselage underside-Check as follows:*1. General condition-Check for skin damage,

such as buckling, splitting, distortion,dents, or fuel leaks.

2. Antennas-Check security and generalcondition.

8-13. NOSE SECTION, AREA 2.a. Nose section-Check as follows:

1. Free air temperature probe-Checkcondition.

2. Avionics door, left side-Check secure.3. Air conditioner exhaust-Check free of

obstruction.4. Wide band data link antenna pod-Check

for cracks and chips.5. Wheel well-Check for signs of leaks,

broken wires and general condition.6. Doors and linkage-Check condition,

security, and alignment.7. Nose gear turning stop-Check condition.

*8. Tire-Check for cuts, bruises, wear,appearance of proper inflation, and wheelcondition.

*9. Shock strut-Check for signs of leakageand 3.0 inches minimum extension.

10. Torque knee-Check condition.11. Shimmy damper and linkage-Check for

security and condition.12. Landing and taxi lights-Check for security

and condition.13. Pitot tubes-Check covers removed,

alignment, security, and free ofobstructions.

14. Radome-Check for cracks and chips.

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TM 55-1510-219-10

CAUTIONDo not move wipers on drywindshield or clean windshieldwith anything other than mildsoap and water.

15. Windshields and wipers-Check windshieldfor cracks and cleanliness and wipers forcontact with glass surface.

16. Air conditioner inlet-Check free ofobstructions.

17. Avionics door, right side-Check secure.8-14. RIGHT WING CENTER SECTION.

a. Right wing center section-Check as follows:1. Deice boot-Check for secure bonding,

cracks, loose patches and generalcondition.

2. Battery access panel-Secure.3. Battery vents-Check free of obstruction.

*4. Auxiliary tank fuel and cap-Check fuellevel visually, condition of seal, and captight and properly installed (locking tabaft).

5. Battery compartment drain-Check free ofobstruction.

6. Battery ram air intake-Check free ofobstruction.

7. INS temperature probe-Check conditionand free of obstructions.

8. Auxiliary tank fuel sump drain-Check forleaks.

9. Heat exchanger outlet and inlet-Check forcracks and free of obstructions.

10. Monopole antenna-Check condition.8-15. RIGHT ENGINE AND PROPELLER.

a. Right engine and propeller-Check as follows:CAUTION

A cold oil check is unreliable. Oilshould be checked within 10minutes after stopping engine. Ifmore than 10 minutes haveelapsed, motor engine for 30

seconds, then recheck. If morethan 10 hours have elapsed, runengine for 2 minutes, thenrecheck. Add oil as required. Donot overfill.*1. Engine oil-Check oil level, oil cap secure

(locking tab aft), and access door locked.2. Engine compartment, left side-Check for

fuel and oil leaks, security of oil cap, doorlocking pins, and general condition.

*3. Left cowl locks-Locked.4. Left exhaust stack-Check for cracks,

security and free of obstructions.*5. Propeller blades and spinner-Check blade

condition, boots, security of spinner, andfree propeller rotation.

*6. Engine air inlets and ice vane-Check freeof obstruction and ice vane retracted.

7. Bypass door-Check condition.*8. Right cowl locks-Locked.9. Right exhaust stack-Check for cracks,

security and free of obstructions.10. Engine compartment, right side-Check

for fuel and oil leaks, ice vane linkage,door locking pins, and general condition.Lock compartment access door.

8-16. RIGHT MAIN LANDING GEAR.a. Right main landing gear-Check as follows:

1. Fuel sump drains (forward)-Check forleaks.

*2. Tires-Check for cuts, bruises, wear,proper inflation and wheel condition.

3. Brake assembly-Check brake lines fordamage or signs of leakage, brake liningsfor wear (0.25 inch maximum) and brakedeice assembly and bleed air hosecondition and security.

*4. Shock strut-Check for signs of leakageand minimum strut extension (5.50inches).

5. Torque knee-Check condition.

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6. Safety switch-Check condition, wire, andsecurity.

7. Fire extinguisher pressure-Checkpressure within limits.

8. Wheel well, doors, and linkage-Check forsigns of leaks, broken wires, security, andgeneral condition.

8-17. RIGHT WING, AREA 3.a. Right wing-Check as follows:

1. Recessed and heated fuel vents-Checkfree of obstructions.

2. Inverter inlet and exhaust louvers-Checkfree of obstructions.

3. GPU access door-Secured.4. Inboard dipole antenna set-Check for

security and cracks at mounting points,bonding secure, free of cuts and cracks.

5. Wing ice light-Check condition.6. Outboard deice boot-Check for secure

bonding, cracks, loose patches, stallstrips, and general condition.

*7. Tiedown-Released.*8. Main tank fuel and cap-Check fuel level

visually, condition of seal, and cap tightand properly installed.

9. Outboard wing fuel vent-Check free ofobstruction.

10. Outboard antenna set-Check for securityand cracks at mounting points, bondingsecure, free of cuts and cracks.

11. Recognition light-Check condition.12. Wing pod, navigation lights and antennas

(2) -Check condition.13. Static wicks-Check security and condition.14. Controls-Check security and condition of

ground adjustable tab.15. Fuel sump drains (3)-Check for leaks.

16. Flaps-Check for full retraction(approximately 0.25 inch play) and skindamage, such as buckling, splitting,distortion, or dents.

17. Chaff dispenser-Check number of chaffsin payload module and for security.

*18. General condition-Check for skin damage,such as buckling, splitting, distortion,dents, or fuel leaks.

8-18. FUSELAGE RIGHT SIDE, AREA 4.a. Fuselage right side-Check as follows:

* 1. General condition-Check for skin damagesuch as buckling, splitting, distortion ordents.

2. Flare/Chaff dispenser-Check number offlares in payload module and for security.

3. Emergency light-Check condition.4. Beacon-Check condition.5. Aft access door-Check secure.6. Oxygen filler door-Check secure.7. Static ports-Check clear of obstructions.8. ASE antennas (2)-Check.9. Emergency locator transmitter ARMED.

10. Emergency locator transmitter antenna-Check condition.

8-19. EMPENNAGE, AREA 5.a. Empennage-Check as follows:

1. Vertical stabilizer, rudder, and trim tab-Check for skin damage, such as buckling,distortion, or dents, and trim tab rig.

2. Antennas-Check condition.3. Deice boots-Check for secure bonding,

cracks, loose patches, and generalcondition.

4. Horizontal stabilizer and elevator-Checkfor skin damage, such as buckling,distortion and dents.

8-10

TM 55-1510-219-10

NOTEAny difference between theindicated position on the trim tabposition indicator and the actualposition of the elevator trim tabsignifies an unairworthy conditionand must be corrected prior to thenext flight of the aircraft.5. Elevator trim tab-Verify "O" (neutral)

position.WARNING

If the possibility of iceaccumulation on the horizontalstabilizer or elevator exists,takeoff will not be attempted.6. Static wicks (16)-Check installed.7. Position and beacon lights-Check

condition.8. Rotating boom dipole antenna-Check

condition and position.9. Wide band data link antenna pod-Check

for cracks and chips.8-20. FUSELAGE, LEFT SIDE, AREA 6.

a. Fuselage-Check as follows:* 1. General condition-Check for skin damage,

such as buckling, distortion, or dents.2. Static ports-Check clear of obstructions.3. ASE antennas (2)-Check.4. Emergency light-Check condition.5. Cabin door-Check door seal and general

condition.6. Fuselage top side-Check general condition

and antennas.*8-21. INTERIOR CHECK.

1. Cargo/loose equipment-Check secure.2. Cabin door-Locked and checked. Insure

that the cabin door is closed and lockedas follows: Check position of safety armand diaphragm plunger (lift door step) andthat each of the six rotary cam locks alignwithin the sight indicators.

3. Cargo door-Locked and checked. Insurethat the cargo door is closed and lockedas follows:

a. Upper handle position-Closed andlocked (the index marks on each ofthe four rotary cam locks must alignwithin the sight indicators).

b. Lower pin latch handle position-Closed and latched (the indicatormust align with stripe on carrierrod).

NOTEThe untapered shoulder of thelatching pins extend past eachattachment lug.

In addition, the following inspection and test shallbe performed prior to the first flight of the day:

c. Open cabin door-Check that the"CABIN/CARGO DOOR"annunciator light is extinguished.

d. Latch cabin door but do not lock-Check that the "CABIN/CARGODOOR" annunciator lightilluminates.

e. Battery switch ON-Check that the"CABIN/CARGO DOOR"annunciator light is still illuminated.

f. Close and lock the cabin door-Check that the "CABIN/CARGODOOR" annunciator light isextinguished.

g. Battery switch-OFF.4. Emergency exit-Check secure and key

removed.5. Mission cooling ducts-Check open and

free of obstructions.6. Flare/Chaff dispenser preflight test-

Completed.7. Crew briefing-As required.

8-22. BEFORE STARTING ENGINES.* 1. Parking brake-Set.

2. Magnetic compass-Check for fluid,heading and current deviation card.

*3. Pedestal controls-Set as follows.

8-11

TM 55-1510-219-10

CAUTIONMovement of power levers intoreverse range while engines areshut down may result in bendingand damage to control linkages.

a. Power levers-IDLE.b. Propeller levers-As required.c. Condition levers-FUEL CUTOFF.d. Flaps-UP.

4. Lower console switches-Set as follows.a. Flare/Chaff dispenser control-SAFE.b. Avionics-As required.c. Rudder boost switch-ON.

5. Gear alternate engage and ratchethandles-Stowed.

6. Free air temperature gage-Check, notecurrent reading.

7. Instrument panel-Check and set asfollows.

a. Pilot's and copilot's course indicatorswitches -As required.

b. Pilot's and copilot's RMI switches-Asrequired.

c. Pilot's and copilot's MIC switch-Asrequired.

d. Pilot's and copilot's compassswitches-As required.

e. Gyro switches-SLAVE.f Flight instruments-Check

instruments for protective glass,warning flags (10) pilot, 5 copilot),static readings, and headingcorrection card.

g. Radar- OFF.h. APR-39 and APR-44-OFF.i. Engine instruments-Check for

protective glass and static readings.8. Prop sync switch-OFF.9. Mission panel switches and circuit

breakers-Set.10. Subpanels-Check and set as follows.

a. Fire protection test switch-OFF.b. Landing, taxi, and recognition lights-

OFF.

c. Gear-Recheck DN.d. Cabin lights-As Required.e. Pilot's static air source-NORMAL.f. Pilot's and copilot's audio control panels-

As required.NOTE

Do not use alternate static sourceduring takeoff and landing exceptin an emergency. Pilot'sinstruments will show a variationin airspeed and altitude.

11. AC and DC GPU's-As required.12. External power advisory lights-As

required.*13. Battery-ON.14. DC power-Check.

*8-23. FIRST ENGINE START (BATTERY START).CAUTION

Mission control switch should beOFF prior to start.

NOTEThe engines must not be starteduntil after the INS is placed intothe NAV mode or OFF as required.

Starting procedures are identical for both engines.When making a battery start, the right engine should bestarted first. When making a ground power unit (GPU)start, the left engine should be started first due to theGPU receptacle being located adjacent to the rightengine. A crew member should monitor the outsideobserver throughout the enginesstart.

1. Avionics master switch-OFF.NOTE

Use the beacons NIGHT positionfor all ground operations,changing to the DAY position onlywhen taking the active runway fortakeoff (when conditions permit).2. Exterior light switches-As required.

8-12

TM 55-1510-219-10

3. Ignition and engine start switch-ON.Propeller should begin to rotate andassociated IGN ON light shouldilluminate. Associated FUEL PRESS lightshould extinguish.

4. Condition lever (after N1 RPM stabilizes,12% minimum)-LOW IDLE.

CAUTIONMonitor TGT to avoid a hot start.If there is a rapid rise in TGT, beprepared to abort the start beforelimits are exceeded. Duringstarting, the maximum allowableTGT is 1000°C for five seconds. Ifthis limit is exceeded, use ABORTSTART procedure anddiscontinue start. Enter the peaktemperature and duration on DAForm 2408-13.5. TGT and N1-Monitor (TGT 1000°C

maximum, N1 52% minimum).6. Oil pressure-Check (60 PSI minimum).7. Ignition and engine start switch-OFF after

TGT stabilized.8. Condition lever-HIGH IDLE. Monitor

TGT as the condition lever is advanced.9. Generator switch-RESET, then ON.

8-24. SECOND ENGINE START (BATTERY START).1. First engine generator load-50% or less.2. Ignition and engine start switch-ON.

Propeller should begin to rotate andassociated IGN on light should illuminate.Associated FUEL PRESS light shouldextinguish.

3. Condition lever (after N1 RPM passes12% minimum)-LOW IDLE.

CAUTIONMonitor TGT to avoid a hot start.If there is a rapid rise in TGT, beprepared to abort the start beforelimits are exceeded. Duringstarting, the maximum allowableTGT is 10000C for five seconds. Ifthis limit is exceeded, use ABORTSTART procedure anddiscontinue start. Enter the peaktemperature and duration on DAForm 2408-13.

4. TGT and N1-Monitor (TGT 1000ICmaximum, N1 52% minimum).

5. Oil pressure-Check (60 PSI minimum).6. Ignition and engine start switch-OFF after

TGT stabilized.7. Battery charge light-ON.8. Second engine generator-RESET, then

ON.9. Inverter switches-ON, check INVERTER

lights OFF.10. Condition levers-As required.

8-25. ABORT START.1. Condition lever-FUEL CUTOFF.2. Ignition and engine start switch-STARTER

ONLY.3. TGT-Monitor for drop in temperature.4. Ignition and engine start switch-OFF.

8-26. ENGINE CLEARING.1. Condition lever-FUEL CUTOFF.2. Ignition and engine start switch-OFF (1

minute minimum).CAUTION

Do not exceed starter limitation of30 seconds ON and 5 minutesOFF for two starting attempts andengine clearing procedure. Allow30 minutes off before additionalstarter operation.3. Ignition and engine start switch-STARTER

ONLY (15 seconds minimum, 30 secondsmaximum).

4. Ignition and engine start switch-OFF.8-27. FIRST ENGINE START (GPU START).

1. INS-As required.CAUTION

Mission control switch should beOFF prior to start.

8-13

TM 55-1510-219-10

NOTEThe engines must not be starteduntil after the INS is placed intothe NAV mode or OFF as required.2. Avionics master switch-As required.

NOTEUse the beacons NIGHT positionfor all ground operations,changing to the DAY position onlywhen taking the active runway fortakeoff (when conditions permit).3. Exterior light switches-As required.4. Ignition and engine start switch-ON.

Propeller should begin to rotate andassociated IGN ON light shouldilluminate. Associated FUEL PRESS lightshould extinguish.

5. Condition lever (after N1 RPM stabilizes,12% minimum)-LOW IDLE.

CAUTIONMonitor TGT to avoid a hot start.If there is a rapid rise in TGT, beprepared to abort the start beforelimits are exceeded. Duringengine start, the maximumallowable TGT is 1000°C for fiveseconds. If this limit is exceeded,use ABORT START procedure anddiscontinue start. Enter the peaktemperature and duration on DAForm 2408-13.6. TGT and N1-Monitor (TGT 1000°C

maximum, NI 52% minimum).7. Oil pressure-Check (60 PSI minimum).8. Ignition and engine start switch-OFF after

TGT stabilized.9. Condition lever-As required. Monitor TGT

as the condition lever is advanced.10. DC GPU disconnect-As required.11. Generator switch (GPU disconnected)-

RESET, then ON.12. Condition lever-HIGH IDLE.

8-28. SECOND ENGINE START (GPU START).1. Ignition and engine start switch-ON.

Propeller should start to rotate and

associated IGN ON light shouldilluminate. Associated FUEL PRESS lightshould extinguish.

2. Condition lever (after N1 RPM passes12% minimum)-LOW IDLE.

CAUTIONMonitor TGT to avoid a hot start.If there is a rapid rise in TGT, beprepared to abort the start beforelimits are exceeded. Duringengine start, the maximumallowable TGT is 1000°C for fiveseconds. If this limit is exceeded,use ABORT START procedure anddiscontinue start. Enter the peaktemperature and duration on DAForm 2408-13.3. TGT and N1-Monitor (TGT 1000°C

maximum, N1 52% minimum).4. Oil pressure-Check (60 PSI minimum).5. Ignition and engine start switch-OFF after

TGT stabilized.6. Propeller levers-FEATHER.7. GPU-Disconnect. (Check aircraft external

power and mission external power lightextinguished).

8. Propellers levers-HIGH RPM.9. Generator switches-RESET, then ON.

10. Aircraft inverter switches-ON, checkINVERTER lights OFF.

11. Condition levers-As required.8-29. BEFORE TAXIING.

*1. Brake deice-As required. To activate thebrake deice system proceed as follows:

a. Bleed air valves-As required.b. Brake deice switch-DEICE. Check

BRAKE DEICE ON light illuminated.c. Condition levers-HIGH IDLE.

CAUTIONVerify airflow is present from aftcockpit eyeball outlets to insuresufficient cooling for missionequipment.*2. Cabin temperature and mode-Set.

8-14

TM 55-1510-219-10

3. AC/DC power-Check.WARNING

Do not operate radar in congestedareas. Injury could result topersonnel in close proximity tooperating radar.

CAUTIONDo not operate the weather radaror data link systems in an areawhere the nearest effectivesurface is 50 feet or less from theantenna reflector. Scanning suchsurfaces within 50 feet of theantenna reflector may damagereceiver crystals.4. Avionics master power switch-ON as

required.5. Mission panel-Set.

a. Mission control-AUTO.b. Data link high voltage-Standby.c. Antenna select-AUTO.d. Antenna steering-AUTO (check

azimuth).e. Antenna override-AUTO rotate.f. Antenna circuit breaker-As required.

6. Electric elevator trim and autopilot/flightdirector operation-Check as required.

7. Avionics-Check as required.8. Flaps-Check as required.9. Altimeters-Check and set.

*8-30. TAXIING.1. Brakes-Check.2. Flight instruments-Check for normal

operation.3. Mission control panel-Set as required.

a. Data link high voltage-ON asrequired.

8-31. ENGINE RUNUP.1. Propeller manual feathering-Check by

pulling propeller levers aft through detentto FEATHER. Check that propeller willfeather, then advance levers to the HIGHRPM position.

2. Autofeather-Check as follows:a. Condition levers-LOW IDLE.b. Autofeather switch-Hold to TEST.c. Power levers-Advance until

AUTOFEATHER lights areilluminated (approximately 22%torque).

d. #1 power lever-Retard.(1) At approximately18% torque-#2

AUTOFEATHER light out.(2) At approximately12% torque-Both

AUTOFEATHER lights out(propeller starts to feather).

e. #1 power lever-Approximately 22%torque.

f. Repeat steps b thru d for #2 engine.3. Overspeed governors-Check as follows:

a. Power levers-Set approximately1950 RPM (both engines).

b. #1 propeller governor test switch-Hold.

c. #1 propeller RPM 1830 to 1910-Check.

d. Repeat steps b and c for #2 engine.e. Power levers-Set 1800 RPM.

4. Primary governors-Check as follows:a. Power 1800 RPM-Set/check.b. Propeller levers aft to detent-Set.c. Propeller RPM 1600-1640-Check.d. Propeller levers to HIGH RPM-Set.

5. Ice vanes-Check as follows:a. Ice vane switches to EXTEND.

Verify torque drop, TGT increase,and illumination of ICE VANE EXTlight-Check.

b. Ice vane switches to RETRACT.Verify return to original torque andTGT, and ICE VANE lightextinguished-Check.

6. Anti-ice and deice systems-Check asfollows:

a. Left pitot switch ON-Check forloadmeter rise, then OFF.

b. Right pitot switch ON-Check forloadmeter rise, then OFF.

8-15

TM 55-1510-219-10

c. Stall warning switch ON-Check forloadmeter rise, then OFF.

d. Fuel vent switch ON-Check forloadmeter rise, then OFF.

e. Windshield anti-ice switchesNORMAL and HI-Check PILOT andCOPILOT (individually) forloadmeter rise, then OFF.

f. Propeller- AUTO (check 14-18amps).

g. Propeller switches-INNER andOUTER (momentarily), check forloadmeter rise.

h. Surface deice switch AUTO-Checkfor a drop in pneumatic pressureand wing deice boots inflation andafter 6 seconds for a second drop inpneumatic pressure. Check manualposition for proper indications.

i. Antenna deice single cycle auto-Check for drop in pneumaticpressure and boots inflated. Checkmanual position for properindications.

j. Radome anti-ice-ON, check forproper indications.

k. Engine inlet lip heat switches-ON,Check for proper indications.

l. Anti-ice and deice systemsswitches-As required.

7. Beacon-As required (DAY or NIGHT).8. Pneumatic pressure-Check as follows:

a. Condition levers-HIGH IDLE.b. Power levers-IDLE.c. Left bleed air valve switch-PNEU &

ENVIRO OFF.d. Pneumatic pressure-12 to 20 PSI-

Check.e. Left bleed air light-Check

illuminated.f. Right bleed air valve switch-PNEU

& ENVIRO OFF.g. Left and right bleed air off and left

and right bleed air fail lights-Checkilluminated.

h. Left bleed air valve switch-OPEN.i. Left bleed air off, and left and right

bleed air fail lights off, andpneumatic pressure-Check (12 to 20PSI).

j. Right bleed air valve switch-OPEN.k. Right bleed air off light-

Extinguished.9. Pressurization system-Check as follows:

a. Cabin door caution lightextinguished-Check.

b. Storm windows closed-Check.c. Bleed air valve switches OPEN-

Check.d. Cabin altitude 500 feet lower than

field pressure altitude-Set.e. Cabin pressurization switch-TEST

(hold).f. Cabin climb gage descending

indication-Check, then releaseTEST switch.

g. Aircraft altitude set to plannedcruise altitude plus 500 feet-Check(if this setting does not result in aCABIN ALT indication of at least500 feet over takeoff field pressurealtitude, adjust as required).

h. Rate control set between 9 and 12o'clock-Check.

10. Windshield anti-ice-As required.NOTE

If windshield anti-ice is neededprior to takeoff, use normalsetting for a minimum of 15minutes prior to selecting hightemperature to provide adequatepreheating and minimize effectsof thermal shock.

8-32. BEFORE TAKEOFF.1. Autofeather switch-ARM.2. Bleed air valves-As required.3. Ice & rain switches-As required.4. Fuel panel-Check fuel quantity and switch

positions.5. Flight and engine instruments-Check for

normal indications.6. Cabin controller-Set.7. Annunciator panels-Check (Note

indications).8. Propeller levers-HIGH RPM.9. Friction locks-Set.

10. Flaps-As required.11. Trim-Set.12. Avionics-Set.13. Flights controls-Check.14. Departure briefing-Complete.

8-33. LINE UP.1. Transponder-As required.2. Engine auto ignition switch- ARM.

8-16 Change 1

TM 55-1510-219-10

3. Power stabilized-Check approximately25% minimum.

4. Condition levers-LOW IDLE.5. Lights-As required (landing, taxi, beacon).6. Mission control panel-Set.

8-34. TAKEOFF.To aid in planning the takeoff and to obtain

maximum aircraft performance, make full use of theinformation affecting takeoff shown in Chapter 7. Thedata shown is achieved by setting brakes, settingTAKEOFF POWER, and then releasing brakes. Whenrunway lengths permit, the normal takeoff may bemodified by starting the takeoff after power has beenstabilized at approximately 25% torque, then applyingpower smoothly so as to attain full power. This willresult in a smoother takeoff but will significantly increasetakeoff distance.8-35. AFTER TAKEOFF.

WARNINGImmediately after takeoff, the pilotflying the aircraft should avoidadjusting controls located on theaft portion of the extendedpedestal to preclude inducingspatial disorientation due tocoriolis illusion.

As cruise climb airspeed is attained, adjust powerto the climb power setting. The copilot then activatesthe YAW DAMP and checks that the cabin ispressurizing. Both pilots check the wings and nacellesfor fuel or oil leaks. The procedural steps after takeoffare as follows:

1. Gear-UP.2. Flaps-UP.3. Landing lights-OFF.

(4.) Windshield anti-ice-As required.NOTE

Turn windshield anti-ice on tonormal when passing 10,000 feetAGL or prior to entering thefreezing level (whichever comesfirst). Leave on until no longerrequired during descent forlanding. High temperature may beselected as required after aminimum warm-up period of 15minutes.

8-36. CUMB.a. Cruise Climb. Cruise climb is performed at a

speed which is the best combination of climb, fuel bum-off, and distance covered. Set propellers at 1900 RPMand torque as required. Adhere to the followingairspeed schedule as closely as possible.

SL to 10,000 feet ...................... 140 KIAS10,000 to 20,000....................... 130 KIAS20,000 to 31,000 feet................ 120 KIAS

b. Climb-Maximum Rate. Maximum rate of climbperformance is obtained by setting propellers at 2,000RPM, torque at 100% (or maximum climb TGT), andmaintaining best rate-of-climb airspeed. Refer toChapter 7 for best rate-of-climb airspeed for specificweights.

1. Climb power-Set.2. Propeller sync-As required.3. Autofeather-As required.4. Yaw damp- As required.5. Cabin pressurization-Check.6. Wings and nacelles-Check.7. ASE-As required.

a. Flare/chaff dispenser safety pin-Asrequired

b. Flare/chaff function selector switch-As required.

c. APR-39-As required.d. APR-44-As required.

8-37. CRUISE.Cruise power settings are entirely dependent upon

the prevailing circumstances and the type of missionbeing flown. Refer to Chapter 7 for airspeed, powersettings, and fuel flow information. The followingprocedures are applicable to all cruise requirements.

1. Power-Set Refer to the cruise powergraphs contained in Chapter 7. Toaccount for ram air temperature increase,it is essential that temperature beobtained at stabilized cruise airspeed.

2. Wings and nacelles-Check.3. Ice & rain switches-As required. Insure

that anti-ice equipment is activated beforeentering icing conditions.

NOTEIce vanes must be extended whenoperating in visible moisture at5°C or less. Visible moisture .ismoisture in any form, clouds, icecrystals, snow, rain, sleet, hail, orany combination of these.4. Auxiliary fuel gages-Monitor. Insure that

fuel is being transferred from auxiliarytanks. (Chapter 2, Section IV.)

Change 1 8-17

TM 55-1510-219-10

5. Altimeters-Check. Verify that altimetersetting complies with transition altituderequirement.

6. Engine instrument indications-Noted.Check all engine instruments for normalindications and record on appropriateforms for use in engine trend monitoring.

7. Recognition lights-As required.8-38. DESCENT.

Descent from cruising altitude should normally bemade by letting down a cruise airspeed with reducedpower.

NOTECABIN pressure CONTROLLERshould be adjusted prior tostarting descent.

a. Descent-Max Rate (Clean). To obtain themaximum rate of descent in clean configuration,perform the following:

1. Power levers-IDLE.2. Propeller levers-HIGH RPM.3. Flaps-UP.4. Gear-UP.5. Airspeed-Vmo.6. Cabin pressurization-Set. Adjust CABIN

CONTROLLER dial as required andadjust RATE control knob so that cabinrate of descent equals one-third aircraftrate of descent.

7. Ice & rain switches-As required.8. Recognition lights-As required.

b. Descent-Max Rate (Landing Configuration). Ifrequired to descend at a low airspeed (e.g., to conserveairspace or in turbulence), approach flaps and landinggear may be extended to increase the rate and angle ofdescent while maintaining the slower airspeed. Toperform the maximum rate of descent in landingconfiguration, perform the following:

1. Power levers-IDLE.2. Propeller levers-HIGH RPM.3. Flap switch-APPROACH.4. Gear switch-DN.5. Airspeed-184 KIAS maximum.6. Cabin pressurization-Set. Adjust CABIN

CONTROLLER dial as required andadjust RATE control knob so that cabindescent rate equals one-third aircraftdescent rate.

7. Ice and rain switches-As required.

8. Recognition lights- As required8-39. DESCENT-ARRIVAL

Perform the following checks prior to the finaldescent for landing.

1. Cabin pressurization-Set. Adjust CABINCONTROLIER dial as required.

2. Ice & rain switches-As required.(3.) Windshield anti-ice-As required.

NOTESet windshield anti-ice to normalor high as required well beforedescent into icing conditions orinto warm moist air to aid indefogging. Turn off windshieldanti-ice when descent iscompleted to lower altitudes andwhen heating is no longerrequired. This will precludepossible wind screen distortions.4. Lights-ON.5. Altimeters-Set to current altimeter setting.6. ASE-As required.

8-40. BEFORE LANDING.1. Prop sync switch-OFF.2. Autofeather switch-ARM.3. Propeller levers-As required.

NOTEDuring approach, propellersshould be set at 1900 RPM toprevent glideslope interference(ILS approach), provide betterpower response during approach,and minimize attitude changewhen advancing propeller leversfor landing.4. Flap switch (below 202 KIAS)-

APPROACH.5. Gear switch (below 184 KIAS)-DN.6. Rotating boom dipole antenna-Check

stowed.7. Landing lights-As required.8. Brake deice-As required.

8-41. LANDING.Performance data charts for landing computations

assume that the runway is paved, level and dry.Additional runway must be allowed when theseconditions are not met Refer to Chapter 7 for landingdata. Do not consider headwind during landingcomputations; however, if landing must be down

8-18 Change 1

TM 55-1510-219-10

wind, include the tailwind in landing distancecomputations. Plan the final approach to arrive at 50feet over the landing area at APPROACH SPEED (Vref)plus 1/2 wind gust speed. Perform the followingprocedures as the aircraft nears the runway.

1. Autopilot and yaw damp-Disengaged.2. Gear down lights-Check.3. Propeller levers-HIGH RPM.

8-42. GO-AROUND.When a go-around is commenced prior to the

LANDING check, use power as required to climb to, ormaintain, the desired altitude and airspeed. If the go-around is started after the LANDING check has beenperformed, apply maximum allowable power andsimultaneously increase pitch attitude to stop thedescent. Retract the landing gear after insuring that theaircraft will not touch the ground. Retract the flaps toAPPROACH, adjusting pitch attitude simultaneously toavoid an altitude loss. Accelerate to best rate-of-climbairspeed (Vy), retracting flaps fully after attaining the Vrefspeed used for the approach. Perform the followingchecks:

1. Power-Maximum allowable.2. Gear- UP.3. Flaps-UP.4. Landing lights-OFF.5. Climb power-Set.6. Yaw damp-As required.

8-43. AFTER LANDING.Complete the following procedures after the

aircraft has cleared the runway:1. Condition levers-As required.2. Engine auto ignition switch-OFF.3. Ice and rain switches-OFF.4. Flaps-UP.5. Avionics-As required.6. Lights-As required.

NOTEReverse should not be used as anormal practice whenmaneuvering or parking theaircraft.7. Mission control panel-Set (DL HV OFF).

8-44. ENGINE SHUTDOWN.CAUTION

To prevent sustained loads onrudder shock links, the aircraftshould be parked with the nosegear centered.1. Brake deice-OFF.2. Parking brake-Set.3. Landing/taxi lights-OFF.4. Overhead floodlight-As required.5. Cabin temperature mode switch-OFF.6. Autofeather switch-OFF.7. Vent and aft vent blower switches-AUTO.8. INS-OFF.9. Mission equipment-OFF, as required.

10. Inverter switches-OFF.11. Battery condition-Check as required.12. Avionics master switch-OFF.13. TGT-Check. TGT must be 660°C or

below for one minute prior to shutdown.14. Propeller levers-FEATHER.

CAUTIONMonitor TGT during shutdown. Ifsustained combustion isobserved, proceed immediately toABORT START procedure.

15. Condition levers-FUEL CUTOFF.WARNING

Do not turn exterior lights off untilpropeller's rotation has stopped.

16. Exterior lights -OFF.17. Master panel lights - OFF.18. Master switch - OFF19. Keylock switch - OFF.20. Oxygen system - As required.

8-19

TM 55-1510-219-10

8-45. BEFORE LEAVING AIRCRAFT.1. Wheels-Chocked.2. Parking brake-As required.

NOTEBrakes should be released afterchocks are in place (rampconditions permitting).3. Flight controls-Locked.4. Fuel pumps-Set.

a. Standby pumps-OFFb. Aux transfer-AUTOc. Crossfeed-CLOSED

5. Emergency exit lock-As required.6. Mode 4-As required.7. Aft cabin light-OFF.8. Door light-OFF.

CAUTIONIf strong winds are anticipatedwhile the aircraft is unattended,the propellers shall be secured toprevent their windmilling withzero engine oil pressure.9. Walk-around inspection-Complete.

Conduct a thorough walk-aroundinspection, checking for damage, fluidleaks, and levels. Check that covers,tiedowns, restraints, safety pins andchocks are installed as required.

10. Aircraft forms-Complete. In addition toestablished requirements for reporting anysystem defects, unusual and excessiveoperation such as hard landings, etc., theflight crew will also make entries on DAForm 2408-13 to indicate when limits inthe Operator's Manual have beenexceeded.

11. Aircraft secured-Check; lock cabin dooras required.

NOTEA cold oil check is unreliable. Oilshould be checked within 10minutes after stopping engine.

Section III. INSTRUMENT FLIGHT

8-46. GENERAL.This aircraft is qualified for operation in instrument

meteorological conditions. Flight handling, stabilitycharacteristics and range are approximately the sameduring instrument flight conditions as when under visualflight conditions.8-47. INSTRUMENT TAKEOFF.

Complete the BEFORE TAKEOFF check.Engage the heading (HDG) mode on the autopilotcomputer/control. (DO NOT ENGAGE AUTOPILOT.)Set heading marker (HDG) to runway heading andadjust pitch bar. Align the aircraft with the runwaycenterline, insuring that nosewheel is straight beforestopping aircraft. Hold brakes and complete the,LINEUP check. Insure that the roll steering bar iscentered. Power application and copilot duties areidentical to those prescribed for a "visual" takeoff. Afterthe brakes are released, initial directional control shouldbe accomplished predominantly with the aid of outsidevisual references. As the takeoff progresses, thecrosscheck should transition from outside references to

the Flight Director and airspeed indicator. The rate oftransition is directly proportional to the rate at which theoutside references deteriorate. Approaching rotationspeed (Vr), the cross-check should be totally committedto the instruments so that erroneous sensory inputs canbe ignored. At rotation speed, establish takeoff attitudeon the Flight Director. Maintain this pitch attitude andwings-level attitude until the aircraft becomes airborne.When both the vertical velocity indicator and altimetershow positive climb indications, retract the landing gear.After the landing gear is retracted, adjust the pitchattitude as required to attain best rate-of-climb airspeed(Vy). Use PITCH-SYNC as required to reposition theFlight Director pitch steering bar. Retract flaps afterobtaining best single-engine rate-of-climb speed (Vyse),and readjust pitch as required. Control the bankattitudes to maintain the desired heading. SupportFlight Director indications throughout the maneuver bycross-checking "raw data" information displayed onsupporting instruments.

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NOTEDue to possible precession error,the pitch steering bar may lowerslightly during acceleration,causing the pitch attitude to appearhigher than actual pitch attitude.To avoid lowering the noseprematurely, crosscheck thevertical velocity and altimeter toinsure proper climb performance.The erection system willautomatically remove the error afterthe acceleration ceases.

8-48. AUTOMATIC APPROACHES.

There are no special preparations required forplacing the aircraft under autopilot control. Refer toChapter 3 for procedures to be followed for automaticapproaches.

NOTEThe ILS localizer and glideslopewarning flags indicate insufficientsignal strength to the receiver.

Section IV. FLIGHT CHARACTERISTICS

8-49. STALLS.

A prestall warning in the form of very lightbuffeting can be felt when a stall is approached. An auralwarning is provided by a warning horn. The warning hornstarts approximately five to ten knots above stall speedwith the aircraft in any configuration.

a. Power-On Stalls. The power-on stall attitude isvery steep and unless this high-pitch attitude ismaintained, the aircraft will generally "settle" or "mush"instead of stall. It is difficult to stall the aircraftinadvertently in any normal maneuver. A light buffetprecedes most stalls, and the first indication ofapproaching stall is generally a decrease in controleffectiveness, accompanied by a "chirping" tone from thestall warning horn. The stall itself is characterized by arolling tendency if the aircraft is allowed to yaw. Theproper use of rudder will prevent the tendency to roll. Aslight pitching tendency will develop if the aircraft is heldin the stall, resulting in the nose dropping sharply, thenpitching up toward the horizon; this cycle is repeated untilrecovery is made. Control is regained very quickly withlittle altitude loss, providing the nose is not loweredexcessively. Begin recovery with forward movement ofthe control wheel and a gradual return to level flight. Theroll tendency caused by yaw is more pronounced inpower-on stalls, as is the pitching tendency; however,both are easily controlled after the initial entry. Power-onstall characteristics are not greatly affected by wing flapposition, except that stalling speed is reduced inproportion to the degree of wing flap extension.

b. Power-Off Stalls. The roll tendency isconsiderably less pronounced in power-off stalls (in anyconfiguration) and is more easily prevented or correctedby adequate rudder and aileron control, respectively. Thenose will generally drop straight through with sometendency to pitch up again if recovery is not made

immediately. With wing flaps down, there is little or noroll tendency and stalling speed is much slower than withwing flaps up. The Stall Speed Chart (Fig. 8-2) showsthe indicated power-off stall speeds with aircraft in variousconfigurations. Altitude loss during a full stall will beapproximately 800 feet.

c. Accelerated Stalls. The aircraft gives noticeablestall warning in the form of buffeting when the stalloccurs. The stall warning and buffet can be demonstratedin turns by applying excessive back pressure on thecontrol wheel.

8-50. SPINS.

Intentional spins are prohibited. If a spin isinadvertently entered use the following recoveryprocedure:

NOTESpin demonstrations have not beenconducted. The recoverytechnique is based on the bestavailable information. The firstthree actions should be as nearlysimultaneously as possible.

1. Power levers IDLE.

2. Apply full rudder opposite the direction of spinrotation.

3. Simultaneously with rudder application, push thecontrol wheel forward and neutralize ailerons.

4. When rotation stops, neutralize rudder.

8-21

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Figure 8-2. Stall Speeds8-22

TM 55-1510-219-10

CAUTIONDo not pull out of the resultingdive too abruptly as this couldcause excessive wing loadsand a possible secondary stall.

5. Pull out of dive by exerting a smooth, steadyback pressure on the control wheel, avoidingan accelerated stall and excessive aircraftstresses.

8-51. DIVING.

Maximum diving airspeed (red line) is 245KIAS or 0. 47 Mach. Flight characteristics are

conventional throughout a dive maneuver; however,caution should be used if rough air is encounteredafter maximum allowable dive speed has beenreached, since it is difficult to reduce speed in diveconfiguration. Dive recovery should be very gentleto avoid excessive aircraft stresses.

8-52. MANEUVERING FLIGHT.

The maximum speed (Va) at which abruptfull control inputs can be applied without exceedingthe design load factor of the aircraft (Va 170 KIAS isshown in Chapter 5). The data is based on 14,200pounds. There are no unusual characteristics underaccelerated flight.

Section V. ADVERSE ENVIRONMENTAL CONDITIONS

8-53. INTRODUCTION.The purpose of this part is to inform the pilot

of the special precautions and procedures to befollowed during the various weather conditions thatmay be encountered in flight. This part is primarilynarrative, only those checklists that cover specificprocedures characteristic of weather operations areincluded. The checklist in Section II provides foradverse environmental operations.

8-54. COLD WEATHER OPERATIONS.Operational difficulties maybe encountered

during extremely cold weather, unless proper stepsare taken prior to or immediately after flight. Allpersonnel should understand and be fully aware ofthe necessary procedures and precautions involved.

a. Preparation For Flight. Accumulations ofsnow, ice, or frost on aircraft surfaces will adverselyaffect takeoff distance, climb performance and stallspeeds to a dangerous degree. Such accumulationsmust be removed before flight. In addition to thenormal exterior checks, following the removal of ice,snow, or frost, inspect wing and empennage surfacesto verify that these remain sufficiently cleared. Also,move all control surfaces to confirm full freedom ofmovement. Assure that tires are not frozen to wheelchocks or to the ground. Use ground heaters, antiicesolution, or brake deice to free frozen tires. Whenheat is applied to release tires, the temperatureshould not exceed 71°C (160°F). Refer to Chapter 2for anti-icing, deicing, and defrosting treatment.

b. Engine Starting. When starting engines onramps covered with ice, propeller levers must be inthe FEATHER position to prevent the tires fromsliding. To prevent exceeding torque limits when

advancing CONDITION levers to HIGH IDLE duringthe starting procedure, place the . power lever inBETA and the propeller lever in HIGH RPM beforeadvancing the condition lever to HIGH IDLE.

c. Warm-Up and Ground Test. Warm-upprocedures and ground test are the same as thoseoutlined in Section II.

d. Taxiing. Whenever possible, taxiing in deepsnow, light weight dry show or slush should beavoided, particularly in colder FAT conditions. If it isnecessary to taxi through snow or slush, do not setthe parking brake when stopped. If possible, do notpark the aircraft in snow or slush deep enough toreach the brake assemblies. Chocks or sandbagsshould be used to prevent the aircraft from rollingwhile parked. Before attempting to taxi, activate thebrake deice system, insuring that the bleed air valvesare OPEN and that the condition levers are in HIGHIDLE. An outside observer should visually checkwheel rotation to insure brake assemblies have beendeiced. The condition levers may be returned toLOW IDLE as soon as the brakes are free of ice.

e. Before Takeoff If icing conditions areexpected, activate all anti-ice systems beforetakeoff, allowing sufficient time for the equipment tobecome effective.

f. Takeoff. Takeoff procedures for coldweather operations are the same as for normaltakeoff. Taking off with temperatures at or belowfreezing, with water, slush or snow on the runway,can cause ice to accumulate on the landing gear andcan throw ice into the wheel well areas. Suchtakeoffs

8-23

TM 55-1510-219-10

shall be made with brake deice on and with the icevanes extended to preclude the possibility of icebuild-up on engine air inlets. Monitor oiltemperatures to insure operation within limits.Before flight into icing conditions, the pilot andcopilot WSHLD ANTI-ICE switches should be set atNORMAL position.

g. During Flight.

(1) After takeoff from a runway coveredwith snow or slush, it may be advisable to leavebrake deice ON to dislodge ice accumulated from thespray of slush or water. Monitor BRAKE DEICEannunciator for automatic termination of systemoperation and then turn the switch OFF. Duringflight, trim tabs and controls should also be exercisedperiodically to prevent freezing. Insure that anti-icingsystems are activated before entering icingconditions. Do not activate the surface deice systemuntil ice has accumulated one-half to one inch. Thepropeller deice system operates effectively as ananti-ice system and it may be operated continuouslyin flight. If propeller imbalance due to ice doesoccur, it may be relieved by increasing RPM briefly,then returning to desired setting.

Ice vanes must be extended whenoperating in visible moisture or when freedom fromvisible moisture cannot be assured, at 5°C FAT orless. Ice vanes are designed as an anti-ice system,not a deice system. After the engine air inlet screensare blocked, lowering the ice vanes will not rectifythe condition. Ice vanes should be retracted at 15°CFAT and above to assure adequate engine oilcooling.

(2) Stalling airspeeds should be expectedto increase when ice has accumulated on the aircraftcausing distortion of the wing airfoil. For the samereason, stall warning devices are not accurate andshould not be relied upon. Keep a comfortablemargin of airspeed above the normal stall airspeed.Maintain a minimum of 140 knots during sustainedicing conditions to prevent ice accumulation onunprotected surfaces of the wing. In the event ofwindshield icing, reduce airspeed to 226 knots orbelow.

h. Descent. Use normal procedures in SectionII. Brake icing should be considered if moisture wasencountered during previous ground operations orinflight in icing conditions with gear extended.

i. Landing. Landing on an icy runway shouldbe attempted only when absolutely necessary andshould not be attempted unless the wind is within 10degrees of runway heading. Application of brakeswithout skidding the tires on ice is very difficult, due

to the sensitive brakes. In order not to impair pilotvisibility, reverse thrust should be used with cautionwhen landing on a runway covered with snow orstanding water. Use procedures in Section II fornormal landing.

j. Engine Shutdown. Use normal proceduresin Section II.

k. Before Leaving the Aircraft. When theaircraft is parked outside on ice or in a fluctuatingfreeze-thaw temperature condition the followingprocedures should be followed in addition to thenormal procedures in Section II. After wheel chocksare in place, release the brakes to prevent freezing.Fill fuel tanks to minimize condensation, remove anyaccumulation of dirt and ice from the landing gearshock struts, and install protective covers to guardagainst possible collection of snow and ice.

8-55. DESERT OPERATION AND HOT WEATHEROPERATION.

Dust, sand, and high temperaturesencountered during desert operation can sharplyreduce the operational life of the aircraft and itsequipment. The abrasive qualities of dust and sandupon turbine blades and moving parts of the aircraftand the destructive effect of heat upon the aircraftinstruments will necessitate hours of maintenance ifbasic preventive measures are not followed. Inflight, the hazards of dust and sand will be difficult toescape, since dust clouds over a desert may befound at altitudes up to 10,000 feet. During hotweather operations, the principle difficultiesencountered are high turbine gas temperatures(TGT) during engine starting, over-heating of brakes,and longer takeoff and landing rolls due to the higherdensity altitudes. In areas where high humidity isencountered, electrical equipment (such ascommunication equipment and instruments) will besubject to malfunction by corrosion, fungi andmoisture absorption by nonmetallic materials.

a. Preparation For Flight. Check the position ofthe aircraft in relation to other aircraft. Propellersand blast can damage closely parked aircraft.Check that the landing gear shock struts are free ofdust and sand. Check instrument panel and generalinterior for dust and sand accumulation. Open mainentrance door and cockpit vent storm windows toventilate the aircraft.

8-24

TM 55-1510-219-10

CAUTION

N1 speeds of 70% or higher may be required tokeep oil temperature within limits.

b. Engine Starting. Use normal procedures in SectionII. Engine starting under conditions of high ambienttemperatures may produce a higher than normal TGTduring the start. The TGT should be closely monitoredwhen the condition lever is moved to the LO IDLE position.If overtemperature tendencies are encountered the condi-tion lever should be moved to IDLE CUTOFF positionperiodically during acceleration of gas generator RPM (N1).Be prepared to abort the start before temperature limitationsare exceeded.

c. Warm-Up Ground Tests. Use normal procedures inSection II. To minimize the possibility of damage to theengines during dusty/sandy conditions, activate ICEVANES if the temperature is below 15°C.

d. Taxiing. Use normal procedures in Section II.When practical, avoid taxing over sandy terrain to mini-mize propeller damage and engine deterioration that resultsfront impingement of sand and gravel. During hot weatheroperation, use minimum braking action to prevent overheat-ing.

e. Takeoff. Use normal procedures in Section II.Avoid taking off in the wake of another aircraft if the run-way surface is sandy or dusty.

f. During Flight. Use normal procedures in Section II.

g. Descent. Use normal procedures in Section II.

h. Landing. Use normal procedures in Section II.

i. Engine Shutdown. Use normal procedures inSection II.

CAUTION

During hot weather, if fuel tanks are completelyfilled, fuel expression may cause overflow,thereby creating a fire hazard.

j. Before Leaving Aircraft. Use normal procedures inSection II. Take extreme care to prevent sand or dust fromentering the fuel and oil system during servicing. Duringhot weather, release the brakes immediately after installingwheel chocks to prevent brake disc warpage.

8-56. TURBULENCE AND THUNDERSTORMOPERATION.

CAUTION

Due to the comparatively light wing loading,control in severe turbulence and thunderstormsis critical. Since turbulence imposes heavyloads on the aircraft structure, make all neces-sary changes in aircraft attitude with the leastamount of control pressures to avoid excessiveloads on the aircraft's structure.

Thunderstorms and areas of severe turbulence should beavoided. If such areas are to be penetrated. it will be neces-sary to counter rapid changes in attitude and accept majorindicated altitude variations. Penetration should be of analtitude which provides adequate maneuvering margins as aloss or gain of several thousand feet of altitude may beexpected. The recommended penetration speed in severeturbulence is 170 KIAS. Pitch attitude and constant powersettings arc vital to proper flight technique. Establishrecommended penetration speed and proper attitude prior toentering turbulent air to minimize most difficulties. Falseindications by the pressure instruments due to barometricpressure variations within the storm make them unreliable.Maintaining a preestablished attitude will result in a fairlyconstant airspeed. Turn cockpit and cabin lights on to mini-mize the blinding effects of lighting. Do not use autopilotaltitude hold. Maintain constant power settings and pitchattitude regardless of airspeed or altitude indications. Con-centrate on maintaining a level attitude by reference to theFlight Director/attitude indicator. Maintain original reading.Make no turns unless absolutely necessary.

Change 1 8-25

TM 55-1510-219-10

8-57. ICE AND RAIN (TYPICAL).

W A R N I N G

While in icing conditions, if there is anunexplained 30% increase of torque neededto maintain airspeed in level flight, acumulative total of two or more inches of iceaccumulation on the wing, an unexplaineddecrease of 15 knots IAS, or an unexplaineddeviation between pilot’s and copilot’sairspeed indicators, the icing environmentshould be exited as soon as practicable. Iceaccumulation on the pitot tube assembliescould cause a complete loss of airspeedindication.

The following conditions indicate a possibleaccumulation of ice on the pitot tube assemblies andunprotected airplane surfaces. If any of these conditionsare observed, the icing environment should be exited assoon as practicable.

(1) Total ice accumulation of two inches or more onthe wing surfaces. Determination of ice thickness can beaccomplished by summing the estimated ice thickness onthe wing prior to each pneumatic boot deice cycle (e.g. fourcycles of minimum recommendation ½-inch accumulation.

(2) A 30 percent increase in torque per enginerequired to maintain a desired airspeed in level flight (notto exceed 85 percent torque) when operating atrecommended holding/loiter speed.

(3) A decrease in indicated airspeed of 15 knotsafter entering the icing condition (not slower than 1.4power off stall speed) if maintaining original power settingin level flight. This can be determined by comparing pre-icing condition entry speed to the indicated speed after asurface and antenna deice cycle is completed.

(4) Any variations from normal indicated airspeedbetween the pilot’s and copilot’s airspeed indicators.

a. Typical Ice. Icing occurs because of supercooledwater vapor such as fog, clouds or rain. The most severeicing occurs on aircraft surfaces in visible moisture orprecipitation with a true outside air temperature between -5°C and +1°C; however, under some circumstances,dangerous icing conditions may be encountered withtemperatures below -10°C. The surface of the aircraftmust be at a temperature of freezing or below before icewill stick to the aircraft. If severe icing conditions areencountered, ascend or descend to altitudes where theseconditions do not prevail. If flight into icing conditions is

unavoidable, proper use of aircraft anti-icing and deicingsystems may minimize the problems encounteredApproximately 15 minutes prior to flight into temperatureconditions which could produce frost or icing conditions,the pilot and co-pilot windshield anti-ice switches shouldhe set at normal or high temperature position (afterpreheating) as necessary to eliminate windshield ice.Stalling airspeeds should be expected to increase when icehas accumulated on the aircraft causing distortion of thewing airfoil. For the same reason, stall warning devices arenot accurate and should not be relied upon. Keep acomfortable margin of airspeed above the normal stallairspeed with ice on the aircraft. Maintain a minimum of140 knots during sustained icing conditions to prevent iceaccumulation on unprotected surfaces of the wing. In theevent of windshield icing, reduce airspeed to 226 knots orbelow.

b. Rain. Rain presents no particular problems otherthan restricted visibility and occasional incorrect airspeedindications.

c. Taxiing. Extreme are must be exercised whentaxiing on ice or slippery runways. Excessive use of eitherbrakes or power may result in an uncontrollable skid.

d. Takeoff. Extreme care must be exercised duringtakeoff from ice or slippery runways. Excessive use ofeither brakes or power may result in an uncontrollable skid.

e. Climb. Keep aircraft attitude as flat as possible andclimb with higher airspeed than usual, so that the lowersurfaces of the aircraft will not be iced by flight at a highangle of attack.

f. Cruising Flight. Prevention of ice formation is farmore effective and satisfactory than attempts to dislodgethe ice after it has formed. If icing conditions areinadvertently encountered, turn on the anti-icing systemsprior to the first sign of ice formation.

Do not operate deicer boots continuously. Allow atleast one-half inch of ice on the boots before activating thedeicer boots to remove the ice. Continued flight in severeicing conditions should not be attempted. If ice forms onthe wing area aft of the deicer boots, climb or descend to analtitude where conditions are less severe.

g. Landing. Extreme care must be exercised whenlanding on ice or slippery runways. Excessive use of eitherbrakes or power may result in an uncontrollable skid. Iceaccumulation on the aircraft will result in higher stallingairspeeds due to the change in aerodynamic characteristicsand increased weight of the aircraft due to ice buildup.Approach and landing airspeeds must be increasedaccordingly.

8-26 Change 3

TM 55-1510-219-10

NOTE

When operating on wet or icy runways, refer tostopping distance factors shown in Chapter 7.

8-57A. ICING (SEVERE).

a. The following weather conditions may be conduciveto severe in-flight icing:

(1) Visible rain at temperatures below zero degreesCelsius ambient air temperature.

(2) Droplets that splash or splatter on impact attemperatures below zero degrees Celsius ambient airtemperature.

b. The following procedures for exiting a severe icingenvironment are applicable to all flight phases from takeoffto landing.

(1) Monitor the ambient air temperature. Whilesevere icing may form at temperatures as cold as -18degrees Celsius, increased vigilance is warranted attemperatures around freezing with visible moisture present.

(2) Upon observing the visual cues specified in thelimitations section of the airplane flight manual (MilitaryOperations Manual) for the identification of severe icingconditions (reference paragraph 5-31B.), accomplish thefollowing:

(a) Immediately request priority handling fromair traffic control to facilitate a route or an altitude changeto exit the severe icing conditions in order to avoidextended exposure to flight conditions more severe thanthose for which the airplane has been certificated.

(b) Avoid abrupt and excessive maneuveringthat may exacerbate control difficulties.

(c) Do not engage the autopilot.

(d) If the autopilot is engaged, hold the controlwheel firmly and disengage the autopilot.

(e) If an unusual roll response oruncommanded roll control movement is observed, reducethe angle-of-attack.

(f) Do not extend flaps during extendedoperation in icing conditions. Operations with flapsextended can result in a reduced angle-of-attack, with thepossibility of ice forming on the upper surface further afton the wing than normal, possibly aft of the protected area.

(g) If the flaps are extended, do not retractthem until the airframe is clear of ice.

(h) Report these weather conditions to airtraffic control.

Change 3 8-26.1

TM 55-1520-219-10

Section VI. CREW DUTIES

8-58. CREW/PASSENGER BRIEFING.

The following guide should be used in accomplishingrequired passenger briefings. Items that do not pertain to aspecific mission may be omitted.

a. Crew introduction.

b. Equipment.

1. Personal, to include ID tags.

2. Professional (medical equipment, etc.).

3. Survival.

c. Flight data.

1. Route.

2. Altitude.

3. Time enroute.

4. Weather.

d. Normal Procedures.

1. Entry and exit of aircraft.

2. Seating and seat position.

3. Seat belts.

4. Movement in aircraft.

5. Internal communications.

6. Security of equipment.

7. Smoking.

8. Oxygen.

9. Refueling.

10. Weapons and prohibited items.

11. Protective masks.

12. Toilet.

e. Emergency Procedures.

1. Emergency Exits.

8-26.2 Change 3

TM 55-1510-219-10

2. Emergency equipment.

3. Emergency landing/ditchingprocedures.

8-59. DEPARTURE BRIEFING.

The following is a guide that should be usedas applicable in accomplishing the required crewbriefing prior to takeoff; however, if the crew hasoperated together previously and the pilot is certainthat the copilot understands all items of the briefing,he may omit the briefing by stating "standardbriefing," when the briefing is called for during theBEFORE TAKEOFF CHECK.

a. ATC Clearance Review.1. Routine.

2. Initial altitude.

b. Departure Procedure Review.1. SID.

2. Noise abatement procedure.

3. VFR departure route.

c. Copilot Duties Review.1. Adjust takeoff power.

2. Monitor engine instruments.

3. Power check at 65 knots.

4. Call out engine malfunctions.

5. Tune/ident all Nav/Com radios.

6. Make all radio calls.

7. Adjust transponder and radar asrequired.

8. Complete flight log during flight (notealtitudes and headings).

9. Note departure time.

d. PPCReview.1. Takeoff power.

2. Vr.

3. Vy (climb to 500' AGL).

4. Vyse

8-60. ARRIVAL BRIEFING.

The following is a guide that should be usedas applicable in accomplishing the required crewbriefing prior to landing; however, if the crew hasoperated together previously and the pilot is certainthat the copilot understands all items of the briefing,he may omit the briefing by stating "standardbriefing," when the briefing is called for during theDESCENT-ARRIVAL CHECK.

a. Weather/Altimeter Setting.b. Airfield/Facilities Review.

1. Field elevation.

2. Runway length.

3. Runway condition.

c. Approach Procedure-Review.

1. Approach plan/profile.

2. Altitude restrictions.

3. Missed approach.

a. Point.

b. Time.

c. Intentions.

4. Decision height or MDA.

5. Lost communications.

d. Back Up Approach/Frequencies.e. Copilot Duties Review.

1. Nav/Com set-up.

2. Monitor altitude and airspeeds.

3. Monitor approach.

4. Callout visual/field in sight.

f. Landing Performance Data Review.1. Approach speed.2. Runway required.

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TM 55-1510-219-10

CHAPTER 9

EMERGENCY PROCEDURES

Section I. AIRCRAFT SYSTEMS

9-1. AIRCRAFT SYSTEMS.This section describes the aircraft systems

emergencies that may reasonably be expected tooccur and presents the procedures to be followed.Emergency procedures are given in checklist formwhen applicable. A condensed version of theseprocedures is in the Operator's and Crewmember'sChecklist, TM 55-1510-219-CL. Emergencyoperations of avionics equipment are covered whenappropriate in Chapter 3, Avionics, and are repeatedin this section only as safety of flight is affected.

9-2. IMMEDIATE ACTION EMERGENCY CHECKS.Immediate action emergency items are

underlined for your reference and shall be committedto memory. During an emergency, the checklist willbe called for to verify the memory steps performedand to assist in completing any additional emergencyprocedures.

NOTEThe urgency of certainemergencies requiresimmediate action by the pilot.The most important singleconsideration is aircraftcontrol. All procedures aresubordinate to thisrequirement.

9-3. DEFINITION OF LANDING TERMS.The term LANDING IMMEDIATELY is

defined as executing a landing without delay. (Theprimary consideration is to assure the survival ofoccupants. ) The term LAND AS SOON ASPOSSIBLE is defined as executing a landing at thenearest suitable landing area without delay. Theterm LAND AS SOON AS PRACTICABLE is definedas executing a landing to the nearest suitable airfield.

9-4. AFTER EMERGENCY ACTION.After a malfunction has occurred,

appropriate emergency actions have been taken, andthe aircraft is on the ground, an entry shall be made

in the remarks section of DA Form 2408-13describing the malfunction.

9-5. EMERGENCY EXITS AND EQUIPMENT.Emergency exits and equipment are shown

in figure 9-1.

9-6. EMERGENCY ENTRANCE.Entry may be made through the cabin

emergency hatch. The hatch may be released bypulling on its flush-mounted pull-out handle,placarded EMERGENCY EXIT-PULL HANDLE TORELEASE. The hatch is of the nonhinged plug typewhich removes completely from the frame when thelatches are released. After the latches are released,the hatch may be pushed in.

9-7. ENGINE MALFUNCTION.a. Flight Characteristics Under Partial Power

Conditions. There are no unusual flightcharacteristics during single-engine operation as longas airspeed is maintained at or above minimumcontrol speed (Vmc) and power-off stall speeds. Thecapability of the aircraft to climb or maintain levelflight depends on configuration, gross weight,altitude, and free air temperature. Performance andcontrol will improve by feathering the propeller of theinoperative engine, retracting the landing gear andflaps, and establishing the appropriate single-enginebest rate-of-climb speed (Vyse). Minimum controlspeed (Vmc) with flaps retracted is approximately 1knot | higher than with flaps at takeoff (40%) position.

b. Engine Malfunction During And AfterTakeoff The action to be taken in the event of anengine malfunction during takeoff depends onwhether or not liftoff speed (V1of) has been attained.If an engine fails immediately after liftoff, manyvariables such as airspeed, runway remaining,aircraft weight, altitude at time of engine failure, andsingle-engine performance must be considered indeciding whether it is safer to land or continue flight.

c. Engine Malfunction Before Liftoff (Abort). Ifan engine fails and the aircraft has not accelerated

9-1

TM 55-1510-219-10

Figure 9-1. Emergency Exits and Equipment

9-2

TM 55-1520-219-10

to recommend liftoff speed (V1of), retard power levers 1. Power - Maximum controllable.immediately to IDLE and stop the aircraft with brakes andreverse thrust. Perform the following: 2. Gear - UP.

3. Flaps - UP.

4. Landing lights - OFF.

5. Brake deice - OFF.

6. Engine cleanup - Perform.

7. Generator load - 100% Maximum.

N O T E

If the autofeather system fails to operate,identify the affected engine, then feather thepropeller of the affected engine.

(3) Engine Malfunction after Liftoff (FlightContinued without Autofeather). Perform the following:

1. Power - Maximum controllable.

N O T E

If airspeed is below Vyse, maintain whateverairspeed has been attained (between V1of andVyse) until sufficient altitude can be obtained totrade off altitude for airspeed to assist inacceleration to Vyse.

2. Dead engine - Identify.

3. POWER lever (dead engine) - IDLE.

4. PROP lever (dead engine) - FEATHER.

5. GEAR-UP.

N O T E

If takeoff was made with flaps extended, insurethat airspeed is above computed approach speed(Vref) before retracting flaps.

6. FLAPS-UP.

7. LANDING LIGHTS - OFF.

8. BRAKE DEICE - OFF.


10. Generator load - 100% Maximum.

1. Power levers - IDLE.

2. Braking as required.

N O T E

If insufficient runway remains for stopping,perform steps 3. through 5.

3. Condition levers - FUEL CUT-OFF.

4. Fire pull handles - Pull.

5. Master switch - OFF.

d. Engine Malfunction after Liftoff. If an engine failsafter becoming airborne, maintain single-engine best rate-of-climb speed (Vyse) or, if airspeed is below Vyse maintainwhatever airspeed is attained between liftoff (V1of) and Vyse

until sufficient altitude is attained to trade altitude forairspeed and accelerate to Vyse.

(1) Engine Malfunction after Liftoff (Abort).Perform the following and land in a wings-level attitude:

1. Power levers - IDLE.

2. Gear - Down.

N O T E

If able to land on remaining runway, use brakesand reverse thrust as required, then performsteps 3. through 5.


4. Fire pull handles - Pull.


(2) Engine Malfunction after Liftoff (FlightContinued). Perform the following:

N O T E

Do not retard the malfunctioning engine powerlever, or turn the autofeather system OFF untilpropeller is completely feathered. To do so willdeactivate the autofeather circuit and preventautomatic feathering.

(9-2.1 blank)9-2.2 Change 3

e. Engine Malfunction During Flight.

1.

2.

3. Dead Engine - IDENTIFY.

4.

5.

6 .

7.

8.

9.

10.

Autopilot/yaw damper - Disengage.

Power - As required.

Power lever (affected engine) - IDLE.

Propeller lever (affected engine) - FEATHER.

Gear - As required.

Flaps - As required.

Power - Set for single engine cruise.

Engine cleanup - Perform.

Generator load - 100% Maximum.

f. Engine Malfunction During Final Approach. If anengine malfunctions during a final approach (afterLANDING CHECK), the propeller should not be manuallyfeathered unless time and altitude permit or conditionsrequire it. Continue approach using the followingprocedure:

1. Power - As required.

2. Gear - DN.

g. Engine Malfunction (Second Engine). If the secondengine fails, do not feather the propeller if an engine restartis to be attempted. Engine restart without starter assist cannot be accomplished with a feathered propeller, and thepropeller will not unfeather without the engine operating.140 KIAS is recommended as the best all-around glidespeed (considering engine restart, distance covered,transition to landing configuration, etc.), although it doesnot necessarily result in the minimum rate of descent.Perform the following procedure if the second engine failsduring cruise flight.

1. Airspeed - 140 KIAS.

2. Power lever - IDLE.

3. Propeller lever - Do not FEATHER.

4. Conduct engine restart procedure.

9-8. ENGINE SHUTDOWN IN FLIGHT.

If it becomes necessary to shut an engine down duringflight, perform the following:


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Change 3 9-3

TM 55-1510-219-10

2. Propeller lever - FEATHER.

3. Condition lever - FUEL CUTOFF.


9-9. ENGINE CLEANUP.

The cleanup procedure to be used after enginemalfunction, shutdown, or an unsuccessful restart isas follows:


2. Engine auto ignition switch - OFF.

3. Autofeather switch - OFF.

4. Generator switch - OFF.

5. Prop sync switch - OFF.

6. Radome heat - OFF.USING STARTER).

9-10. ENGINE RESTART DURING FLIGHT (us-ING STARTER).

Engine restarts may be attempted at all alti-tudes. If a restart is attempted, perform the follow-ing:

1. Cabin temperature mode switch - OFF.

2. Electrical load - Reduce to minimum.

3. Fire pull handle - In.




7. TGT (operative engine) - 700°C or less.

8. Ignition and engine start switch - ON.

9. Condition lever - LOW IDLE.

N O T E

If a rise in TGT does not occur within 10seconds after moving the condition leverto LOW IDLE, abort the start.

10. TGT - 1000°C, 5 seconds maximum.

NOTE

If N1 is below 12%, starting temperaturestend to be higher than normal. To pre-clude overtemperature (1000°C or above)during engine acceleration to idle speed,periodically move the condition lever intoFUEL CUTOFF position as necessary.

11.

12.

13.

14.

15.

Oil pressure - Check.

Ignition and engine start switch - OFF.

Generator switch - RESET, then ON.

Engine cleanup - Perform if engine restartunsuccessful.

Cabin temperature mode switch - As re-quired.

16. Electrical equipment - As required.

17. Auto ignition switch - ARMED.

18. Propellers - Synchronized.


9-11. ENGINE RESTART DURING FLIGHT (NOT

A restart without starter assist may be accom-plished provided airspeed is at or above 140 KIASaltitude is below 20,000 feet, and the propeller is notfeathered. If altitude permits, diving the aircraft willincrease N1 and assist in restart. N1 required for air-start should be at or above 9%. If a start is attempt-ed, perform following:

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

Cabin temperature mode switch - OFF.

Electrical load - Reduce to minimum.

Generator switch (affected engine) - OFF.

Fire pull handle - Check in.

Power lever - IDLE.

Propeller lever - HIGH RPM.

Condition lever - FUEL CUTOFF.

Airspeed 140 KIAS minimum - Check.

Altitude below 20,000 feet - Check.

Engine auto ignition switch - ARM.

Condition lever - LOW IDLE.

TGT - 1000°C 5 seconds maximum.

Oil pressure - Check.

Generator switch - RESET then ON.

Engine cleanup - Perform if engine restartunsuccessful.

16. Cabin temperature mode switch - As re-quired.

17. Electrical equipment - As required.

18. Auto ignition switch - ARMED.

19. Propellers - Synchronized.

9-4


9-12. MAXIMUM GLIDE.

In the event of failure of both engines, maximumgliding distance can be obtained by feathering bothpropellers to reduce propeller drag and by maintaining theappropriate airspeed with the gear and flaps up. Figure 9-2gives the approximated gliding distances in relation toaltitude.

9-13. LANDING WITH TWO ENGINES IN-OPERATIVE.

Maintain best glide speed (figure 9-2). If sufficientaltitude remains after reaching a suitable landing area, acircular pattern will provide best observation of surfaceconditions, wind velocity and direction. When thecondition of the terrain has been noted and the landing areaselected, set up a rectangular pattern. ExtendingAPPROACH flaps and landing gear early in the patternwill give an indication of glide performance sooner, andwill allow more time to make adjustments for the addeddrag. Fly the base leg as necessary to control Point oftouch-down. Plan to overshoot rather than undershoot,then use flaps as necessary to arrive at the selected landingpoint.

Keep in mind that, with both propellers feathered, thenormal tendency is to overshoot due to less drag. In theevent a positive gear-down indication cannot bedetermined, prepare for a gear-up landing. Also, unless thesurface of the landing area is hard and smooth, the landingshould be made with the landing gear up. If landing onrough terrain, land in a slightly tail-low attitude to keepnacelles from possibly digging in. If possible, land withflaps fully extended.

9-14. LOW OIL PRESSURE.

In the event of a low oil pressure indication, perform theprocedures below, as applicable:

1. Oil pressure below 105 PSI below 21,000 feet or 85PSI at 21,000 feet and above, torque - 49%maximum.

2. Oil pressure below 60 PSI - Perform engineshutdown, or land as soon as practicable usingminimum power to ensure safe arrival.

9 - 1 5 . C H I P D E T E C T O R W A R N I N GANNUNCIATOR ILLUMINATED.

If the L CHIP DET or R CHIP DET warningannunciator illuminates, and safe single-engine flight canbe maintained, perform the following:

2. Land as soon as practicable.

1. Perform engine shutdown.

9-16. DUCT OVERTEMP CAUTION LIGHTILLUMINATED.

Insure that the cabin floor outlets are open andunobstructed, then perform the following steps in sequenceuntil the light is extinguished. After completion of steps 1.through 4., if the light does not extinguish, allowapproximately 30 seconds after each adjustment for thesystem temperature to stabilize. The overtemperaturecondition is considered corrected at any point during theprocedure that the light goes out.

1. Cabin air control - In.

2. Cabin temperature mode switch - AUTO.

3. Cabin temperature rheostat - Full decrease.

4. Vent blower switch - HI.

5. Cabin temperature mode switch - MAN HEAT.

6. Manual temperature switch - DECREASE (hold).

7. Left bleed air valve switch - ENVIRO OFF.

8. Light still illuminated (30 seconds) - Left bleed airvalve switch - OPEN.

9. Right bleed air valve switch - ENVIRO OFF.

10. Light still illuminated (30 seconds) - Right bleed airvalve switch - OPEN.

N O T E

If the overtemperature light has not extinguishedafter completing the above procedure, thewarning system has malfunctioned.

9-17. ICE VANE FAILURE.

Ice vane failure is indicated by VANE FAIL cautionlight illumination. If an ice vane fails to operateelectrically, perform the following:

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Figure 9-2. Maximum Glide Distance

9-6

After the ice vanes have been manuallyextended, they may be mechanically re-tracted. No electrical extension or retrac-tion shall be attempted as damage to theelectric actuator may result. Linkage inthe nacelle area must be reset prior to op-eration of the electric system. Do not re-set ice vane control circuit breaker.

1. Airspeed - 160 KIAS or below.

2. Ice vane control circuit breaker - Pull.

3. Ice vane - Operate manually. masks - 100% and on.

4. Airspeed - Resume normal airspeed.

DO NOT RETRACT ice vanes electrical-ly after manual extension.

9-18. ENGINE BLEED AIR SYSTEM MALFUNC-TION.

a. Bleed Air Failure Light Illuminated. Steadyillumination of the warning light in flight indicatesa possible ruptured bleed air line aft of the enginefirewall. The light will remain illuminated for the re-mainder of flight. Perform the following:

NOTE

BLEED AIR FAIL lights may momentari-ly illuminate during simultaneous surfacedeice and brake deice operation at lowN1. speeds.

1. Brake deice switch - OFF.

2. TGT and torque - Monitor (note read-ings).

3. Bleed air valve switch - PNEU & EN-VIRO OFF.

NOTE

Brake deice on the affected side, and rud-der boost , will not be available withBLEED AIR VALVE switch in PNEU &ENVIRO OFF.

4. Cabin pressurization - Check.

b. Excessive Differential Pressure. If cabin dif-ferential pressure exceeds 6.1 PSI, perform the fol-lowing:

1.

2.

3.

4.

5.

6.

7.

Cabin controller - Select higher setting.

If condition persists: Left bleed airvalve switch - ENVIRO OFF flight illu-minated).

If condition still persists: Right bleedair valve switch - ENVIRO OFF (lightilluminated).

If condition still persists - Descend im-mediately.

If unable to descend: Crew oxygen

If unable to descend: CABIN PRESSswitch - DUMP.

Bleed air valve switches - OPEN, ifcabin heating is required.

9-19. LOSS OF PRESSURIZATION (ABOVE 10,000 FEET).

If cabin pressurization is lost when operatingabove 10,000 feet or the ALTITUDE warning lightilluminates, perform the following:

1. Crew oxygen masks - 100% and on.

9-20. CABIN DOOR CAUTION LIGHT ILLUMINAT-ED.

Remain clear of cabin door and perform the fol-lowing:

1. Bleed air valve switches - ENVIRO OFF.

2. Descend below 14,000 feet as soon as practi-cable.

3. Oxygen - As required.

9-21. SINGLE-ENGINE DESCENT/ARRIVAL.

N O T E

Approximately 85% N1 is required tomaintain pressurization schedule.

Perform the following procedure prior to the fi-nal descent for landing.

1. Cabin controller - Set.

2. Seat belts and harnesses - Secure.

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9 - 7

TM 55-1510-219-10

3.

4.titude.

5.

6.

Ice and rain switches - As required.

Altimeters - Set.

Recognition lights - ON.

Arrival briefing - Complete.

9-22. SINGLE-ENGINE BEFORE LANDING.

1. Propeller lever - As required.

N O T E

During approach, propeller should be setat 1900 RPM to prevent glideslope inter-ference (ILS approach), provide betterpower response during approach, and tominimize attitude change when advancingpropeller levers for landing.

2.

3.

4.

5.

6.

Flaps - APPROACH.

Gear - DN.

Landing lights - As required.

Yaw damp - OFF.

Brake deice - OFF.

9-23. SINGLE-ENGINE LANDING CHECK. 9-26. FIRE.

Perform the following procedure during final ap-proach to runway.

1. Autopilot/yaw damp - Disengaged.

2. Gear lights - Check.

3. Propeller lever (operative engine) - HIGHRPM.

N O T E

To insure constant reversing characteris-tics. the propeller control must be in theHIGH RPM position.

9-24. SINGLE-ENGINE GO-AROUND.lowing:

Once flaps are fully extended, a single-engine go-around may not be possible

when close to grouud under conditions ofhigh gross weights and/or high density al-

1. Power - Maximum controllable.

2 . G e a r - U P .

3. Flaps - As required.

4. Landing lights - OFF.


6. Yaw damp - As required.

9-25. PROPELLER FAILURE (OVER 2000 RPM).

If an overspeed condition occurs that cannot becontrolled with the propeller lever, or by reduciugpower, perform the following:

1. Power lever (affected engine) - IDLE.


3. Condition lever - As required.4. Engine cleanup - As required,

The safety of aircraft occupants is the primaryconsideration when a fire occurs; therefore, it is im-perative that every effort be made by the flight crewto put the fire out. On the ground it is essential thatthe engines be shut down, crew evacuated, and firefighting begun immediately. If the aircraft is air-borne when a fire occurs, the most important singleaction that can be taken by the pilot is to land safelyas soon as possible.

a. Engine Fire. The following procedures shallbe performed in case of engine fire:

(1) Engine/nacel le f ire during s tart orground operations. If engine/nacelle fire is identifiedduring start or ground operation, perform the fol-

1. Propeller levers - FEATHER.


3. Fire pull handle - Pull.

9-8 Change 3

C A U T I O N

If fire extinguisher has been used to extinguishan engine fire, do not attempt to restart untilmaintenance personnel have inspected theaircraft and released it for flight.

4. Push to extinguish switch - Push.


(2) Engine fire in flight (fire pull handle lightilluminated). If an engine fire is suspected in flight,perform the following:

1.

2.

3.

Power lever - IDLE.

Fire pull handle out - Advance power.

Fire pull handle light still illuminated -Perform engine fire in flight procedures(identified).

N O T E

Flight into the sun at high aircraft pitch attitudemay actuate the fire warning system. Loweringthe nose and/or changing headings will confirma warning system failure caused by sun rays.

(3) Engine fire in flight (identified). If an enginefire is confirmed in flight, perform the following:

Due to the possibilities of fire warning systemmalfunctions, the fire should be visuallyidentified before the engine is secured and theextinguisher actuated.



3. Condition lever - FUEL CUT-OFF.


5. Fire extinguisher - Actuate as required.


b. Fuselage Fire. If a fuselage fire occurs, perform thefollowing:

The extinguisher agent (Bromochlorodifluoro-methane) in the fire extinguisher can producetoxic effects if inhaled.

1. Fight the fire.

2. Land as soon as possible.

c. Wing Fire. There is little that can be done to controla wing fire except to shut off fuel and electrical systemsthat may be contributing to the fire, or which couldaggravate it. Diving and slipping the aircraft away fromthe burning wing may help. If a wing fire occurs, performthe following:

1. Perform engine shutdown on affected side.


d. Electrical Fire. Upon noticing the existence orindications of an electrical fire, turn off all affectedelectrical circuits, if known. If electrical fire source isunknown, perform the following:

1. Crew oxygen - 100%.

2. Master switch - OFF (Visual Conditions Only).

3. All nonessential electrical equipment - OFF.

N O T E

With loss of DC electrical power, the aircraftwill depressurize. All electrical instruments,with the exception of the Prop RPM, N1 RPM,and TGT gages will be inoperative.

4. Battery switch - ON.

5. Generator switches (individually) - RESET, thenON.

6. Circuit breakers - Check for indication ofdefective circuit.

CAUTION

As each electrical switch is returned to ON, noteloadmeter reading and check for evidence offire.

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Change 3 9-9

TM 55-1520-219-10

7. Essential electrical equipment - On (individuallyuntil fire source is isolated).


e. Smoke and Fume Elimination. To eliminate smokeand fumes from the aircraft, perform the following:

1.

2.

3.

4.

5.

6.

Crew oxygen - 100% and on.

Bleed air valve switches - ENVIRO OFF.

Vent blower switch - AUTO.

Aft vent blower switch - OFF.

Cabin temperature mode switch - OFF.

If smoke and fumes are not eliminated, CABINPRESS switch - DUMP.

N O T E

Opening storm window (after depressurizing)will facilitate smoke and fume removal.

7. Engine oil pressure - Monitor.

9-27. FUEL SYSTEM.

a. Fuel Press Warning Light Illuminated. Illuminationof the #1 or #2 FUEL PRESS warning light usuallyindicates failure of the respective engine-driven boostpump. Perform the following:

1. Standby pump switch - ON.

2. Fuel pressure light out - Check.

3. Fuel pressure light still on - Record unboostedtime.

b. No Fuel Transfer Caution Light Illuminated.Illumination of the #1 or #2 NO FUEL XFR light with fuelremaining in the respective auxiliary fuel tank indicates afailure of that automatic fuel transfer system. Proceed asfollows:

1. AUX TRANSFER switch (affected side)OVERRIDE.

2. Auxiliary fuel quantity - Monitor.

3. AUX TRANSFER switch (after respectiveauxiliary fuel has completely transferred) -AUTO.

c. Nacelle Fuel Leak. If nacelle fuel leaks are evident,perform the following:

1. Perform engine shutdown.



d. Fuel Crossfeed. Fuel crossfeed is normally usedonly during single-engine operation. The fuel from thedead engine side may be used to supply the live engine byrouting the fuel through the crossfeed system. Duringextended flights, this method of fuel usage will provide amore balanced lateral load condition in the aircraft. Forfuel crossfeed. use the following procedure:

1. AUX TRANSFER switches - AUTO.

N O T E

With the FIRE PULL handle pulled, the fuel inthe auxiliary tank for that side will not beavailable (usable) for crossfeed.

2. Standby pumps - OFF.

3. Crossfeed switch - As required.

4. Fuel crossfeed light illuminated - Check.

N O T E

With the FIRE PULL handle pulled, the FUELpress light will remain illuminated on the sidesupplying fuel.

5. Fuel pressure light extinguished - Check.

6. Fuel quantity - Monitor.

9-10 Change 3

TM 55-1510-219-10

e. NAC LOW Light Illuminated. Illumination of the #1or #2 NAC LOW caution light indicates that the affectedtank has 20 minutes remaining at sea level, maximumcruise power consumption rate. Proceed as follows:

W A R N I N G

Failure of the fuel tank venting system willprevent the fuel in the wing tanks from gravityfeeding into the nacelle tank. Fuel vent systemfailure may be indicated by illumination of the#1 or #2 NAC LOW caution light with greaterthan 20 minutes of usable fuel indicated in themain tank fuel system. The total usable fuelremaining in the main fuel supply system withthe LOW FUEL caution light illuminated maybe as little as 114 pounds, regardless of the totalfuel quantity indicated. Continued flight mayresult in engine flameout due to fuel starvation.

1. Twenty minutes fuel remaining - Confirm.


9-28. ELECTRICAL SYSTEM EMERGENCIES.

a. DC GEN Light Illuminated. Illumination of a #1 or#2 DC GEN caution light indicates failure.

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Change 3 9-10.1/(9-10.2 blank)

TM 55-1510-219-10

of a generator or one of its associated circuits(generator control unit). If one generator systembecomes inoperative, all nonessential electricalequipment should be used judiciously to avoidoverloading the remaining generator. The use ofaccessories which create a very high drain should beavoided. If both generators are shut off due to eithergenerator system failure or engine failure, allnonessential equipment should be turned off topreserve battery power for extending the landinggear and wing flaps. When a DC GEN lightilluminates, perform the following:

1. Generator switch-OFF, RESET, then ON.2. Generator switch (no reset)-OFF.3. Mission control switch-OVERRIDE.4. Operating loadmeter-100% maximum.

b. Both DC GEN Lights Illuminated.

1. All nonessential equipment-OFF.2. Land as soon as practicable.

c. Excessive Loadmeter Indication (Over100%). If either loadmeter indicates over 100%,perform the following:

1. Battery switch-OFF (monitor loadmeter).2. Loadmeter over 100%-Nonessential

electrical equipment OFF.3. Loadmeter under 100%-BATT switch ON.

d. Inverter Light Illuminated. Illumination of the#1 or #2 AIRCRAFT INVERTER caution lightindicates failure of the affected inverter. Wheneither inverter fails, the total aircraft load isautomatically switched to the remaining inverter.When a #1 or #2 AIRCRAFT INVERTER lightilluminates, perform the following:

1. Affected AIRCRAFT INVERTER switch-OFF.

e. INST AC Light Illuminated. Illumination ofthe INST AC warning light indicates that 26 VACpower is not available. All items connected to the 26VAC bus will be inoperative. Under these conditions,power must be controlled by indications of the N1and TGT gages. Perform the following:

1. N1 and TGT indications-Check.

2. Other engine instruments-Monitor.

f: Circuit Breaker Tripped. If the circuit breakeris for a nonessential item, do not reset in flight. If thecircuit breaker is for an essential item, the circuitbreaker may be reset once. If a bus feeder circuitbreaker (on the overhead circuit breaker panel) trips,a short is indicated. Do not reset in flight. If a circuitbreaker trips, perform as follows:

1. BUS FEEDER breaker tripped-Do not reset.2. Nonessential circuit-Do not reset.3. Essential circuit-Reset once.

g. BATTERY CHARGE Light Illuminated.

1. Battery volt-amp meter Check.

a. Amp reading above 7 amps anddecreasing Monitor.

b. Amp reading above 7 amps andincreasing Battery switch OFF.

2. Battery switch ON (for landing prior togear/flap extension).

9-29. EMERGENCY DESCENT.

Emergency descent is a maximum effort inwhich damage to the aircraft must be consideredsecondary to getting the aircraft down. The followingprocedure assumes the structural integrity of theaircraft and smooth flight conditions. If structuralintegrity is in doubt, limit speed as much as possible,reduce rate of descent if necessary, and avoid highmaneuvering loads. For emergency descent,perform the following:

NOTEWindshield defogging may berequired.

1. Power lever-IDLE.2. Propeller lever-HIGH RPM.3. Flap lever-APPROACH.4. Gear-DN.5. Airspeed-184 KIAS maximum.

9-11

TM 55-1510-219-10

9-30. LANDING EMERGENCIES.

WARNINGStructural damage may existafter landing with brake, tire, orlanding gear malfunctions.Under no circumstances shallan attempt be made to inspectthe aircraft until jacks havebeen installed.

a. Landing Gear Unsafe Indication. Should oneor more of the landing gear fail to indicate a safecondition, the following steps should be taken beforeproceeding manually to extend the gear.

1. LANDING GEAR RELAY circuitbreaker Check in.

2. Gear lights Check.3. Gear handle DN.4. Manual gear extension As required.

NOTEIf gear continues to indicateunsafe, attempt to verifyposition visually by otheraircraft.

b. Landing Gear Emergency Extension.

CAUTIONAfter an emergency landinggear extension has been made,do not stow the gear ratchethandle or move any landinggear controls or reset anyswitches or circuit breakersuntil the cause of themalfunction has beencorrected.

1. Airspeed 130 KIAS.2. LANDING GEAR RELAY circuit

breaker Out.3. Gear handle DN.4. Landing gear alternate engage handle

Lift and turn clockwise to the stop.5. Alternate landing gear extension

handle Pump.6. Gear lights illuminated Check.

NOTEAfter a practice manualextension, the alternate handle

may be stowed and the landinggear retracted electrically.Rotate the alternate engagehandle counterclockwise andpush it down. Stow thehandle, push in the LANDINGGEAR RELAY circuit breakeron the overhead circuit breakerpanel and retract the gear inthe normal manner with thelanding gear switch.

c. Gear-up Landing (All Gear Up or Unlocked).Due to decreased drag with the gear up, thetendency will be to overshoot the approach. Thecenter-of-gravity with the gear retracted is aft of themain wheels. This condition will allow the aircraft tobe landed with the gear retracted and should result ina minimum amount of structural damage to theaircraft, providing the wings are kept level. It isrecommended that the fuel load be reduced and thelanding made with flaps fully extended on a hardsurface runway. Landing on soft ground or dirt is notrecommended as sod has a tendency to roll up intochunks, damaging the underside of the aircraft'sstructure. When fuel load has been reduced,prepare for a gear-up landing as follows:

1. Crew emergency briefing Complete.2. Loose equipment Stowed.3. Bleed air valves ENVIRO OFF.4. Cabin pressure switch DUMP.5. Cabin emergency hatch Remove and

stow.6. Seat belts and harnesses Secured.7. Landing gear alternate engage handle

Disengaged.8. Alternate landing gear extension

handle Stowed.9. Gear relay circuit breaker In.

10. Gear handle UP.11. Nonessential electrical equipment

OFF.12. Flaps As required (DOWN for landing).

NOTEFly a normal approach totouchdown. After landing,accomplish the following:

13. Condition levers FUEL CUTOFF.14. Fire pull handles Pull.15. Master switch OFF.

9-12

TM 55-1510-219-10

d. Landing With Nose Gear Unsafe. If the LDGGEAR CONTROL warning light is illuminated andthe nose GEAR DOWN LIGHT shows an unsafecondition, the nose gear is probably not locked down,and the gear position should be checked visually byanother aircraft, if possible. If all attempts to lock thenose gear fail, a landing should be made with themain gear down and locked. Hold the nose off therunway as long as possible and do not use brakes.Use the following procedures:


stow.6. Seat belts and harnesses Secured.7. Nonessential electrical equipment

OFF.


8. Condition levers FUEL CUTOFF.9. Fire pull handle Pull.

10. Master switch OFF.

e. Landing With One Main Gear Unsafe. If onemain landing gear fails to extend, retract the othergear and make a gear-up landing. If all efforts toretract the extended gear fail, land the aircraft on ahard runway surface, touching down on the sameedge of the runway as the extended gear. Roll onthe down and locked gear, holding the opposite wingup and the nose gear straight as long as possible. Ifthe gear has extended, but is unsafe, apply brakeslightly on the unsafe side to assist in locking thegear. If the gear has not extended or does not lock,allow the wing to lower slowly to the runway. Use thefollowing procedures:


stow.6. Seat belts and harnesses Secured.7. Nonessential electrical equipment

OFF.8. Touchdown On safe main gear first.


9. Condition levers FUEL CUTOFF.10. Fire pull handle Pull.11. Master switch OFF.

f Landing With Flat Tire(s). If aware that amain gear tire(s) is flat, a landing close to the edge ofthe runway opposite the flat tire will help avoidveering off the runway. If the nose wheel tire is flat,use minimum braking.

9-31. CRACKED WINDSHIELD.a. External Crack. If an external windshield

crack is noted, no action is required in flight.

NOTEHeating elements may beinoperative in areas of crack.

b. Internal Crack. If an internal crack occurs,perform the following:

1. Descend Below 25,000 feet.

2. Cabin Pressure Reset pressuredifferential to 4 PSI or less within 10minutes.

9-32. CRACKED CABIN WINDOW

If crack in a single ply of the external cabinwindow occurs, unpressurized flight may becontinued. Proceed as follows:

1. Oxygen As required.2. Cabin pressurization Depressurize.3. Descend As required.

NOTEIf both plys of the externalcabin window have developedcracks, the aircraft shall not beflown once landed, withoutproper ferry flightauthorization.

9-13

TM 55-1510-219-10

9-33. DITCHING.

If a decision to ditch is made, immediatelyalert all crewmembers to prepare for ditching. Planthe approach into the wind if the wind is high and theseas are heavy. If the swells are heavy but the windis light, land parallel to the swells. Set up a minimumrate descent (power on or off, as the situationdictates airspeed 110-120 KIAS). Do not try to flareas in a normal landing, as it is very difficult to judgealtitude over water, particularly in a slick sea.Leveling off too high may cause a nose low "drop in,"while having the tail too low on impact may result inthe aircraft pitching forward and "digging in. " Expectmore than one impact shock and several skipsbefore the final hard shock. There may be nothingbut spray visible for several seconds while theaircraft is decelerating. To prevent cartwheeling, it isimportant that the wings be level when the aircrafthits the water. After the aircraft is at rest, superviseevacuation of passengers and exit the aircraft asquickly as possible. In a planned ditching, the liferaft and first-aid kits should be secured close to thecabin emergency hatch for easy access whenevacuating; however, do not remove the raft from itscarrying case inside the aircraft. After exiting theaircraft, keep the raft away from any damagedsurfaces which might tear or puncture the fabric.The length of time that the aircraft will float dependson the fuel level and the extent of aircraft damagecaused by the ditching. Refer to Figure 9-3 for bodypositions during ditching, and Table 9-1 for personnelprocedures. Figure 9-4 shows wind swellinformation. Perform the following procedures:

WARNINGDo not unstrap from the seatuntil all motion stops. Thepossibility of injury anddisorientation requires thatevacuation not be attempteduntil the aircraft comes to acomplete stop.

1. Radio calls/transponder-As required.2. Crew emergency briefing-As required.3. Bleed air valves-ENVIRO OFF.4. Cabin pressurization switch-DUMP.5. Cabin emergency hatch-Remove and stow.6. Seat belts and harnesses-Secured.

7. Gear-UP.8. Flaps-DOWN.9. Nonessential electrical equipment-OFF.

10. Approach-Normal, power on.11. Emergency lights-As required. Ditching

9-34. FLIGHT CONTROLS MALFUNCTION.

Use the following procedures, as applicable,for flight control malfunctions.

a. Unscheduled Rudder Boost Activation.Rudder boost operation without a large variation ofpower between engines indicates a failure of thesystem. Perform the following:

1. Rudder boost-OFF.

NOTEThe rudder boost system maynot operate when the brakedeice system is in use.Availability of the rudder boostsystem will be restored tonormal when the BRAKEDEICE switch is turned OFF.

IF CONDITION PERSISTS:

2. Bleed air valve switches-PNEU &ENVIRO-OFF.

3. Rudder trim-Adjust.

b. Unscheduled Electric Elevator Trim. In theevent of unscheduled electric elevator trim, performthe following:

1. ELEV TRIM switch OFF.2. ELEC TRIM circuit breaker OUT.

9-35. BAILOUT.

When the decision has been made toabandon the aircraft in flight. the pilot will give thewarning signal. Exit from the aircraft will be throughthe main entrance door, and in the departuresequence using the exit routes as indicated in Figure9-1. Proceed as follows if bailout becomesnecessary:

1. Notify copilot to prepare to bail out.2. Distress message Transmit.3. Voice security ZEROIZE.

9-14

TM 55-1510-219-10

Table 9-1. Personnel Procedures During Ditching

4. Transponder-7700.

5. Flaps-DOWN.

6. Airspeed-100 KIAS.

7. Trim-As required.

8. Autopilot-Engage.

9. Cabin pressure switch-DUMP.

10. Radio cord, oxygen hose, harnesses andseat belt-Disconnect.

11. Parachute-Attach to harness.

12. Cabin door-Open.

13. Abandon the aircraft.

9-15

TM 55-1510-219-10

BRACE POSITIONS

Figure 9-3. Emergency Body Positions

Figure 9-4. Wind Swell Ditch Heading Evaluation

9-16

TM 55-1510-219-10

APPENDIX A

REFERENCES

Reference information for the subject material contained in this manual can be found in the followingpublications.

AR 70-50 Designating and Naming Defense Equipment, Rockets, and Guided MissilesAR 95-1 Army Aviation - General Provisions and Flight RegulationsAR 95-3 Weight and Balance - Army AircraftAR 380-40 Safeguarding COMSEC InformationAR 385-40 Accident Reporting and RecordsAR 700-26 Aircraft Designation SystemFAR Part 91 General Operating and Flight RulesFM 1-230 Meteorology for Army AviatorsTB 55-9150-200-24 Engine and Transmission Oils, Fuels, and Additives for Army AircraftTB AVN 23-13 Anti-icing, Deicing and Defrosting Procedures for Parked AircraftTB MED 501 Noise and Conservation of Hearing

TM 9-1095-206-13&P Operator's Aviation Unit Maintenance and Aviation IntermediateMaintenance Manual (Including Repair Parts and Special Tools List) toDispenser, General Purpose Aircraft: M-130

(C) TM 11-5825-252-15 Operator, Organizational, DS, GS, and Deport Maintenance Manual: RC-12D Aircraft Mission Equipment, (V)

TM 11-5841-291-12 Operator and Organizational Maintenance Manual, Radar Warning System,AN/APR-44(V) 1

TM 11-5841-283-20 Organizational Maintenance Manual for Detection Set, Radar Signal AN/APR-39(V) 1.

TM 11-6140-203-14-2 Operator's Organizational, Direct Support, General Support and DepotMaintenance Manual Including Repair Parts and Special Tools List: AircraftNickel-Cadmium Batteries

TM 11-6940-214-12 Operator and Organizational Maintenance Manual, Simulator, Radar Signal,SM-756/APR-44(V)

TM 55-410 Aircraft Maintenance, Servicing and Ground Handling Under ExtremeEnvironmental Conditions

TM 55-1500-314-25 Handling, Storage, and Disposal of Army Aircraft Components ContainingRadioactive Materials

TM 55-1500-204-25/1 General Aircraft Maintenance Manual

TM 750-244-1-5 Procedures for the Destruction of Aircraft and Associated Equipment toPrevent Enemy Use

A-1/(A-2 blank)

TM 55-1510-219-10APPENDIX B

ABBREVIATIONS AND TERMSFor the purpose of this manual, the following abbreviations and terms apply. See appropriate technicalmanuals for additional terms and abbreviations.AIRSPEED TERMINOLOGY.

CAS Calibrated airspeed is indicated airspeed corrected for position and instrument er-ror.

FT/MIN Feet per minute.

GS Ground speed, though not an airspeed, is directly calculable from true airspeed ifthe true wind speed and direction are known.

IAS Indicated airspeed is the speed as shown on the airspeed indicator and assumes noerror.

KT Knots.

TAS True airspeed is calibrated airspeed corrected for temperature, pressure, and com-pressibility effects.

Va Maneuvering speed is the maximum speed at which application of full availableaerodynamic control will not overstress the aircraft.

Vf Design flap speed is the highest speed permissible at which wing flaps may be actu-ated.

Vfe Maximum flap extended speed is the highest speed permissible with wing flaps ina prescribed extended position.

Vle Maximum landing gear extended speed is the maximum speed at which an aircraftcan be safely flown with the landing gear extended.

Vlo Maximum landing gear operating speed is the maximum speed at which the landinggear can be safely extended or retracted.

Vlof Lift off speed (takeoff airspeed).

Vmca The minimum flight speed at which the aircraft is directionally controllable as deter-mined in accordance with Federal Aviation Regulations. Aircraft Certification con-ditions include one engine becoming inoperative and windmilling; a 5° bank to-wards the operative engine; takeoff power on operative engine; landing gear up;flaps up; and most rearward CG. This speed has been demonstrated to provide sat-isfactory control above power off stall speed (which varies with weight, configura-tion, and flight attitude).

Vmo Maximum operating limit speed.

Vne Never exceed speed.

Vr Rotation speed.

Vs Power off stalling speed or the minimum steady flight speed at which the aircraftis controllable.

Vso Stalling speed or the minimum steady flight speed in the landing configuration.

Vsse The safe one-engine inoperative speed selected to provide a reasonable marginagainst the occurrence of an unintentional stall when making intentional enginecuts.

B-1

TM 55-1510-219-10

Vx Best angle of climb speed.Vxse Best single-engine angle of climb speed.Vy Best rate of climb speed.Vyse The best single engine rate of climb speed.

METEOROLOGICAL TERMINOLOGY.

Altimeter Setting Barometric pressure corrected to sea level.°C Degrees Celsius.°F Degrees Fahrenheit.FAT Free air temperature is the free air static temperature, obtained either from ground

meteorological sources or from inflight temperature indications adjusted for com-pressibility effects.

Indicated Pressure The number actually read from an altimeter when the barometric scale (KollsmanAltitude window) has been set to 29.92 inches of mercury (1013 millibars).

ISA International standard atmosphere in which:a. The air is a dry perfect gas;b. The temperature at sea level is 59 degrees Fahrenheit, 15 degrees Celsius;c. The pressure at sea level is 29.92 inches Hg;d. The temperature gradient from sea level to the altitude at which the temperatureis -69.7 degrees Fahrenheit is -0.003566 Fahrenheit per foot and zero above that al-titude.

Pressure Altitude Indicated pressure altitude corrected for altimeter error(press alt)

SL Sea level.

Wind The wind velocities recorded as variables on the charts of this manual are to be un-derstood as the headwind or tailwind components of the actual winds at 50 feetabove runway surface (tower winds).

Beta Range The region of the power lever control which is aft of the idle stop and forward ofreversing range where blade pitch angle can be changed without a change of gas generator RPM.

Cruise Climb Is the maximum power approved for normal climb. This power is torque or temper-ature (ITT) limited.

High Idle Obtained by placing the condition lever in the HIGH IDLE position.

HP Horsepower.

Low Idle Obtained by placing the condition lever in the LO IDLE position.

Maximum Cruise Is the highest power rating for cruise and is not time limited.Power

Maximum Power The maximum power available from an engine for use during an emergency opera-tion.

Normal Rated Climb The maximum power available from an engine for continuous normal climb opera-Power tions.

Normal Rated Power The maximum power available from an engine for continuous operation in cruise(with lower ITT limit than normal rated climb power).

B-2

TM 55-1510-219-10

Reverse Thrust Obtained by lifting the power levers and moving them aft of the beta range.

RPM Revolutions per minute.

Takeoff Power The maximum power available from an engine for takeoff, limited to periods of fiveminutes duration.

CONTROL AND INSTRUMENT TERMINOLOGY.Condition Lever The fuel shut-off lever actuates a valve in the fuel control unit which controls the(Fuel Shut-off Lever) flow of fuel at the fuel control outlet and regulates the idle range from LO to HIGH.

Interstage Turbine Eight probes wired in parallel indicate the temperature between the compressor andTemperature (ITT) power turbines.

N1 Tachometer (Gas The tachometer registers the RPM of the gas generator with 100% representing a gasGenerator RPM) generator speed of 37,500 RPM.

Power Lever (Gas This lever serves to modulate engine power from full reverse thrust to takeoff. TheGenerator N1 RPM) position for idle represents the lowest recommended level of power for flight opera-

tion.Propeller Control This lever requests the control to maintain RPM at a selected value and, in the max-Lever (N2 RPM) imum decrease RPM position, feathers the propeller.

Propeller Governor This governor will maintain the selected propeller speed requested by the propellercontrol lever.

Torquemeter The torquemeter system determines the shaft output torque. Torque values are ob-tained by tapping into two outlets on the reduction gear case and recording the dif-ferential pressure from the outlets.

GRAPH AND TABULAR TERMINOLOGY.AGL Above ground level.

Best Angle of Climb The best angle-of-climb speed is the airspeed which delivers the greatest gain of alti-tude in the shortest possible horizontal distance with gear and flaps up.

Best Rate of Climb The best rate-of-climb speed is the airspeed which delivers the greatest gain of alti-tude in the shortest possible time with gear and flaps up.

Clean Configuration Gear and flaps up regardless of mission antenna installation.

Demonstrated The maximum 90° crosswind component for which adequate control of the aircraftCrosswind during takeoff and landing was actually demonstrated during certification tests.

Gradient The ratio of the change in height to the horizontal distance, usually expressed inpercent.

Landing Weight The weight of the aircraft at landing touchdown.

Maximum Zero Fuel Any weight above the value given must be loaded as fuel.Weight

MEA Minimum enroute altitude.

Obstacle Clearance Obstacle clearance climb speed is a speed near Vx and Vy, 1.1 times power off stallClimb Speed speed, or 1.2 times minimum single-engine stall-speed, whichever is higher.

Ramp Weight The gross weight of the aircraft before engine start. Included is the takeoff weightplus a fuel allowance for start, taxi, run up and takeoff ground roll to liftoff.

B-3

TM 55-1510-219-10

Route Segment A part of a route. Each end of that part is identified by:a. A geographic location; orb. A point at which a definite radio fix can be established.

Service Ceiling The altitude at which the minimum rate of climb of 100 feet per minute can be at-tained for existing aircraft weight.

Takeoff Weight The weight of the aircraft at liftoff from the runway.

WEIGHT AND BALANCE TERMINOLOGY.

Arm The distance from the center of gravity of an object to a line about which momentsare to be computed.

Approved Loading Those combinations of aircraft weight and center of gravity which define the limitsEnvelope beyond which loading is not approved.

Basic Empty Weight The aircraft weight with unusable fuel, full oil, and full operating fluids.

Center-of-Gravity A point at which the weight of an object may be considered concentrated for weightand balance purposes.

CG Limits CG limits are the extremes of movement which the CG can have without makingthe aircraft unsafe to fly. The CG of the loaded aircraft must be within these limitsat takeoff, in the air, and on landing.

Datum A vertical plane perpendicular to the aircraft longitudinal axis from which fore andaft (usually aft) measurements are made for weight and balance purposes.

Engine Oil That portion of the engine oil which can be drained from the engine.

Empty Weight The aircraft weight with fixed ballast, unusable fuel, engine oil, engine coolant, hy-draulic fluid, and in other respects as required by applicable regulatory standards.

Landing Weight The weight of the aircraft at landing touchdown.

Maximum Weight The largest weight allowed by design, structural, performance or other limitations.

Moment A measure of the rotational tendency of a weight, about a specified line, mathemati-cally equal to the product of the weight and the arm.

Standard Weights corresponding to the aircraft as offered with seating and interior, avionics,accessories, fixed ballast and other equipment specified by the manufacturer ascomposing a standard aircraft.

Station The longitudinal distance from some point to the zero datum or zero fuselage sta-tion.

Takeoff Weight The weight of the aircraft at liftoff.

Unusable Fuel The fuel remaining after consumption of usable fuel.

Usable Fuel That portion of the total fuel which is available for consumption as determined inaccordance with applicable regulatory standards.

Useful Load The difference between the aircraft ramp weight and basic empty weight.

B-4

TM 55-1510-219-10

MISCELLANEOUS ABBREVIATIONS

Deg Degrees LB PoundsDN Down MAX MaximumFT Foot or feet MH MegahertzFT LB Foot-pounds MIN MinimumGAL Gallons NAUT NauticalHR Hours NM Nautical mileskH Kilohertz PSI Pounds per square inchRC Rate of climb

B-5/(B-6 blank)

TM 55-1510-219-10INDEX

Paragraph, Figure,Subject Table Number


A BAccelerate-Go Flight Path Example ........... 7-8, F7-1 Backup VOW AN/ARC-164 ................. 3-8Accelerate-Stop (Flaps 0%) ...................... 7-6 Baggage Moments .............................. T6-1Accelerometer .......................................... 2-82 Bailout ................................................ 9-35AC Power Supply ...................................... 2-73 Bank and Pitch Limits ......................... 5-28Action Codes and Recommended Action . . T3-4 Before Exterior Check ......................... 8-6ADF Control Panel (DF-203) ..................... F3-16 Before Landing .................................... 8-40After Emergency Action ............................. 9-4 Before Leaving Aircraft ........................ 8-45After Landing ............................................ 8-43 Before Starting Engines ...................... 8-22After Takeoff ............................................. 8-35 Before Takeoff .................................... 8-32Aileron High Torque Test Switch Before Taxiing ..................................... 8-29and Annunciator ....................................... F3-19 Brake Deice Limitations ....................... 5-11Air Conditioning System ........................... 2-68 Brake Deice System ............................ 2-57Air Induction Systems - General ............... 2-18Aircraft Compartments and Stations . ........ 6-3, F6-1 CAircraft Designation System ...................... 1-11Aircraft Systems Emergencies .................. 9-1 Cabin and Cargo Doors........................ F2-7Airspeed Indicators ................................... 2-78 Cabin Door Caution Light .................... 2-10Airspeed Limitations ................................. 5-20 Cabin Door Caution Light Illuminated .. 9-20Altitude Limitations ................................... 5-29 Cabin Door Limitations ........................ 5-39Ammunition............................................... 4-5 Cabin Pressure Limits ......................... 5-34Annunciator Panels.................................... T2-6 Center of Gravity Limitations ............... 5-18,6-11Antenna Deice System .............................. 2-53 Center of Gravity Limits (Landing.........Anti-Icing, Deicing and Defrosting .............. Gear Down) ........................................ T6-4, T6-5Protection ................................................. 2-96 Center of Gravity Moments ................. T6-3Appendix A, References............................. 1-4 Chaff/Flare Dispenser Control Panel ... F4-2Appendix B, Abbreviations and Terms........ 1-5 Chart C - Basic Weight and Balance ....Application of External Power ................... 2-97 Record DD Form 365-3 ....................... 6-7Approved Fuels ........................................ T2-8 Charts and Forms ................................ 6-5Army Aviation Safety Program .................. 1-7 Checklist ............................................. 8-4Arrival Briefing .......................................... 8-60 Chip Detector Warning Light Illuminated . 9-15Audio Control Panels ................................ 3-6, F3-1 Cigarette Lighters and Ash Trays ........ 2-63Autoignition System .................................. 2-31 Class .................................................. 6-2Automatic Approaches .............................. 8-48 Climb .................................................. 8-36Automatic Direction Finder (DF-203) ......... 3-25 Cockpit ............................................... F2-9Automatic Flight Control System ............... 3-27 Cold Weather Operations .................... 8-54Autopilot/Flight Director Annunciator.......... Comments Pertinent To the Use of.......Panel ........................................................ F3-20 Performance Graphs ........................... 7-13Autopilot Limitations ................................. 5-9 Communications - Description ............ 3-5Autopilot Mode Selector Panel ................... Condition Levers ................................. 2-23(614E-42A) ............................................... F3-18 Conditions - Performance .................... 7-2Autopilot Pitch-Turn Control Panel ............ F3-21 Control Locks ..................................... 2-17Auxiliary Fuel Tank Mechanical Gage ....... F2-13 Control Wheels ................................... 2-317 F2-16Avionics Equipment Configuration ............ 3-2 Copilot's Altimeter ............................... 2-80

Copilot's Gyro Horizon Indicator .......... 3-21, F3-14

INDEX-1

TM 55-1510-219-10



C ECopilot's Horizontal Situation Engine Clearing 8-26Indicator (331A-6P) .................................. F3-11 Engine Compartment Cooling .................. 2-17Cracked Cabin Window ............................ 9-32 Engine Fire Detection System .................. 2-25Cracked Windshield ................................. 9-31 Engine Fire Extinguisher Gage Pressure .. T2-1Crew Briefings .......................................... 8-2, 8-58 Engine Fire Extinguisher System .............. 2-26Crossfeed Fuel Flow ................................ F2-15 Engine Fuel Control System ..................... 2-21Cruise ....................................................... 8-37 Engine Ice Protection Systems ................. 2-20Cylinder Capacity vs Pressure and Engine Ignition System ............................ 2-30Temperature ............................................ F2-20 Engine Instruments .................................. 2-33

Engine Limitations ................................... 5-14D Engine Malfunction ................................... 9-7

Engine Operating Limitations ................... T5-1DC Electrical System ............................... F2-22 Engine Restart During Flight .................... 9-10, 9-11DC Power Supply ..................................... 2-72 Engine Runup .......................................... 8-31Definition of Landing Terms ..................... 9-3 Engines - Description ............................... 2-16Defrosting/Defogging System ................... 2-51 Engine Shutdown ..................................... 8-44Departure Briefing .................................... 8-59 Engine Shutdown In Flight ....................... 9-8Descent ................................................... 8-38 Engine Starter-Generators ....................... 2-32Descent-Arrival ........................................ 8-39 Entrance and Exit Provisions ................... 2-9Desert Operation and Hot Weather Environmental Controls ........................... 2-70Operation ................................................. 8-55 Environmental System ............................. F2-21Destruction of Army Materiel To Prevent Exceeding Operational Limits ................... 5-3Enemy Use .............................................. 1-8 Exhaust and Propeller Danger Areas ........ F2-5Dimensions .............................................. 2-3 Exhaust Danger Area ............................... 2-6Ditching ................................................... 9-33, T9-1 Explanation of Change Symbols .............. 1-10Diving ...................................................... 8-51 Extent of Coverage .................................. 6-1Draining Moisture From Fuel System ....... 2-89 Exterior Inspection ................................... F8-1Dual NAV I - NAV 2 Control Panel ............ F3-15 Exterior Inspection - Center Section,Duct Overtemp Caution Light Illuminated . 9-16 Left Side .................................................. 8-11

Exterior Inspection - Empennage, Area 5 . 8-19E Exterior Inspection - Fuselage Left Side,

Area 6 ..................................................... 8-20Electrical Power Supply and Distribution Exterior Inspection - Fuselage Right Side,System - Description ................................ 2-71 Area 6 ..................................................... 8-18Electrical System Emergencies ................ 9-28 Exterior Inspection - Fuselage Underside . 8-12Electric Toilet ........................................... 2-64 Exterior Inspection - Left Engine andEmergency Body Positions ....................... F9-3 Propeller .................................................. 8-10Emergency Descent ................................. 9-29 Exterior Inspection - Left MainEmergency Entrance ................................ 9-6 Landing Gear ........................................... 8-9Emergency Equipment - Description ......... 2-13 Exterior Inspection - Left Wing, Area 1...... 8-8Emergency Exits and Equipment............... 9-5, F9-1 Exterior Inspection - Nose Section, Area 2 8-13Emergency Lighting ................................. 2-76 Exterior Inspection - Right Engine andEmergency Locator Transmitter (ELT) ...... 3-15, F3-8 Propeller .................................................. 8-15Engine Bleed Air System Malfunction ....... 9-18 Exterior Inspection - Right MainEngine Chip Detection System ................. 2-29 Landing Gear ........................................... 8-16Engine Cleanup ........................................ 9-9 Exterior Inspection - Right Wing, Area 3 ... 8-17

Exterior Inspection - Right Wing,Center Section.......................................... 8-14

INDEX-2

TM 55-1510-219-10

Subject Paragraph, Figure,Table Number

E

Exterior Lighting . . . . . . . . . . . . . . . . . . . . . . . . 2-74, F2-27

F

Feathering Provisions . . . . . . . . . . . . . . . . . . . . . . . . . 2-44Ferry Chair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-36Ferry Fuel System 2-36Filling Fuel Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-88Fire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-26First Aid Kits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-14First Engine Start (Battery Start)

. . . . . . . . . . . . . . . . . . . . . . . .2-40

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8-23First Engine Start (GPU Start) 8-27Flight Controls - DescriptionFlight Controls Lock . . . . . . . . . . . . . . . . . . . . . . . . . . .

2-37

Flight Controls Malfunction . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-34Flight EnvelopeFlight Planning

. . . . . . . . . . . . . . . . . . . . . . . . . .

Flight Under IMC

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

F5-27-9

5-31Foreign Object Damage Control 2-19Forms and Records . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-9Free Air Temperature (FAT) Gage . . . . . . . . . . . . . . 2-83Friction Lock Knobs . . . . . . . . . . . . . . . . . . . . . . . . 2-24Fuel and Oil Data . . . . . . . . . . . . . . . . . . . . . . . . . . 6-10Fuel Moments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T6-2Fuel Handling Precautions . . . . . . . . . . . . . . . . . . . . . 2-87Fuel Load . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-9Fuel, Lubricants, Specifications

and Capacities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T2-7Fuel Management Panel & Auxiliary

Tank Mechanical Gage . . . . . . . . . . . . . . . . . . . . . . F2-13Fuel Quantity Data . . . . . . . . . . . . . . . . . . . . . . . . . . . T2-2Fuel Sample

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

8-7

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Fuel Sump Drain Locations . . . . . . . . . . . . . . . . . . . . . . . T2-3Fuel Supply System. . . . . . . . . . . . . . . . . . . . . . . . . . . 2-34Fuel System Anti-Icing

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2-58Fuel System Emergencies 9-27Fuel System LimitsFuel System Management

. . . . . . . . . . . . . . . . . . . . . . .5-102-35

Fuel System Schematic F2-12Fuel Types . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-90

G

General Exterior Arrangement . . . . . . . . . . . . . . . . . . F2-1General Interior Arrangement . . . . . . . . . . . . . . . . . . . F2-2 Landing Gear SystemGenerator Limits . . . . . . . . . . . . . . . . . . . . . . . 5-17, T5-2

Subject Paragraph, Figure,Table Number

G

Go Around . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-42Gravity Feed Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-14Ground Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-99Ground Turning Radius . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

2-4, F2-4Gyromagnetic Compass Systems

. . . . . . . . . . . . . . . . .

3-22

H

HF Command Set (718 U-5) . . . . . . . . . . . . . . 3-14, F3-7Hand-Operated Fire Extinguisher

. . . . . . . . . . . . . . . . . . . . . . .

2-15Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-67Horizontal Reference Indicator

. . . . . . . . . . . . .

3-19, F3-12Horizontal Situation Indicators . . . . . . . 3-18, F3-10, F3-11

I

INS Control Display Unit (C-IV-E) F3-23INS Mode Selector Unit (C-IV-E)

. . . . . . . . . . . . . . . . .F3-22

Ice and Rain (Typical). . . . . . . . . . . . . . . . . . . . . . . .

8-57Icing Limitations (Typical)

. . . . . . . . . . . . . . . . . . . . . . . . .5-31A

Icing Limitations (Severe) 5-31BIcing (Severe) 8-57AIce Vane Failure

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

Immediate Action Emergency Checks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

............................

9-179-2

Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-6Inertial Navigation System . . . . . . . . . . . . . . . . . . . . . 3-28Inflating Tires

. . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-94

Installation of Protective Covers 2-100Instrument Landing System Limits . . . . . . . . . . . . . . . 5-35Instrument Marking Color Codes . . . . . . . . . . . . . . . 5-6Instrument Markings . . . . . . . . . . . . . . . . . . . . 5-5, F5-1Instrument Panel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-28Instrument Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-47Intentional Engine Cut Speed. . . . . . . . . . . . . . . . . . 5-37Interior Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-21Interior Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-75Introduction to Performance . . . . . . . . . . . . . . . . . . . . . 7-1

L

Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-41Landing Emergencies 9-30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Landing Gear Extension Speed .. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .Landing Gear Retraction Speed5-225-23

. . . . . . . . . . . . . . . . . . . . . . . . . . 2-7Landing Information . . . . . . . . . . . . . . . . . . . . . . . . . 7-12Landing With Two Engines Inoperative

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . 9-13Line Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-33

Change 3 INDEX-3

TM 55-1510-219-10

Loading Procedure . . . . . . . . . . . . . . . . . . 6-13Load Planning . . . . . . . . . . . . . . . . . . . . . 6-12Loss of Pressurization (Above 10,000 Feet) . . . 9-19Low Oil Pressure . . . . . . . . . . . . . . . . . . . .9-14

M

M-130 Flare and Chaff DispensingSystem . . . . . . . . . . . . . . . . . . . . . . . . 4-3

Malfunction Codes . . . . . . . . . . . . . . . . . . T3-5Malfunction Codes Check . . . . . . . . . . . . . T3-2Malfunction Indications and Procedures . . . T3-3Maneuvering Flight . . . . . . . . . . . . . . . . . 8-52Maneuvers . . . . . . . . . . . . . . . . . . . . . . . 5-27Marker Beacon . . . . . . . . . . . . . . . . . . . . 3-24Maximum Allowable Airspeed . . . . . . . . . . 5-21Maximum Design Maneuvering Speed . . . . . 5-26Maximum Design Sink Rate 5-40Maximum Glide

. . . . . . . . . . .

. . . .

. . . . . . . . . . . . . . . 9-12Maximum Glide Distance . . . . . . . . . . . . . F9-2Maximum Weights . . . . . . . . . . . . . . . . . . . 2-5Microphones, Switches and Jacks . . . . . . . . . 3-4Minimum Crew Requirements . . . . . . . . . . . 5-4Minimum Single-Engine Control Airspeed 5-25Miscellaneous Instruments

. . . .. . . . . . . . . . . . . . . . 2-85

Mission Annunciators . . . . . . . . . . . . . . . . T4-1Mission Avionics Operating Instructions 4-1Mission Control Panel . . . . . . . . . . . . 4-2, F4-1Mission Control Switches . . . . . . . . . . . . . T4-2Mission Equipment Power System . . . . . . F2-25Mission Planning . . . . . . . . . . . . . . . . . . . . 8-1Mooring . . . . . . . . . . . . . . . . . . . . . . . . 2-101Mooring the Aircraft . . . . . . . . . . . . . . . . F2-32

N

Navigation - Description . . . . . . . . . . . . . . . 3-16

O

Oil Supply System . . . . . . . . . . . . . . . . . . 2-27Operating Procedures and Maneuvers . . . . . . 8-3Overhead Circuit Breaker Panel . . . . . . . . F2-26Overhead Control Panel . . . . . . . . . . . . . F2-18Overtemperature and Overspeed

Limitations . . . . . . . . . . . . . . . . . . . . . . 5-15

Paragraph, FigureSubject Table Number Subject

Paragraph, Figure,Table Number

L O

Oxygen Duration In Minutes . . . . . . . . . . . T2-5Oxygen Flow Planning Rates vs Altitude . . . T2-4Oxygen Requirements . . . . . . . . . . . . . . . . 5-33Oxygen System . . . . . . . . . . . . . . . 2-61, F2-19Oxygen System Servicing . . . . . . . . . . . . . F2-30

P

Parking . . . . . . . . . . . . . . . . . . . . . . . . . 2-102Parking Brake . . . . . . . . . . . . . . . . . . . . . . 2-8Parking, Covers, Ground Handling,

and Towing Equipment . . . . . . . . . . . . F2-31Pedestal . . . . . . . . . . . . . . . . . . . . . . . . F2-11Personnel Procedures During Ditching . . . . . T9-1Pilot and Copilot Seats . . . . . . . . . . . . . . . F2-8Pilot’s Encoding Altimeter . . . . 2-79, 3-31, F3-26Pilot’s Horizontal Situation

Indicator (331A-8G) . . . . . . . . . . . . . . F3-10Pilot’s Turn and Slip Indicator . . . . . 3-20, F3-13Pitot and Stall Warning Heat System . . . . . 2-55Pitot and Static System . . . . . . . . . . . . . . . 2-56Pitot Heat Limitations . . . . . . . . . . . . . . . 5-12Pneumatic Surface Deice System Limits . . . 5-13Placard Items . . . . . . . . . . . . . . . . . . . . . 1-13Power Definitions for Engine Operations . . . 5-16Power Levers . . . . . . . . . . . . . . . . . . . . . . 2-22Pressure Altitude . . . . . . . . . . . . . . . . . . . . 7-3Pressurization System . . . . . . . . . . . . . . . . 2-60Principal Dimensions . . . . . . . . . . . . . . . . F2-3Propeller Electrothermal Anti-Ice System 2-54. . .Propeller Failure (Over 2080 RPM) 9-25Propeller Governors

. . . . . . .. . . . . . . . . . . . . . . . . . . . 2-45

Propeller Levers . . . . . . . . . . . . . . . . . . . . 2-48Propeller Limitations . . . . . . . . . . . . . . . . . 5-7Propeller Reversing . . . . . . . . . . . . . . . . . . 2-49Propeller Synchrophaser . . . . . . . . . . . . . . 2-47Propeller Tachometers . . . . . . . . . . . . . . . . 2 -50Propeller Test Switches . . . . . . . . . . . . . . . 2-46Propellen - Description . . . . . . . . . . . . . . . 2-43PT6A-41 Engine . . . . . . . . . . . . . . . . . . . F2-10

R

Radar Signal Detecting Set Control PanelAPR-39(V)1 . . . . . . . . . . . . . . . . . . . . . F 4 - 3

Radar Signal Detecting Set ControlPanel APR-39(V)2) . . . . . . . . . . . . . . . . F4-4

INDEX-4

TM 55-1510-219-10


R

Radar Signal Detecting Set Indicator ..................F4-5Radar Signal Detecting Sets .....................4-6Radar Warning Receiver ControlPanel (AN/APR-44( )(V3)) .........................4-7, F4-6Radio Control Panel ..................................3-7, F3-2Radio Magnetic Indicators (RMI) ...............3-17, F3-9Radome Anti-ice Operation .......................5-38Recommended Fluid Dilution Chart ...........T2-10Relief Tube ...............................................2-66Required Equipment Listing .......................5-41, T5-3Reserve Fuel ............................................7-10Responsibility - Weight and Balance .........6-6Rudder System .........................................2-39

S

Seats ........................................................2-12Second Engine Start (Battery Start) ..........8-24Second Engine Start (GPU Start) ..............8-28Securing Loads..........................................6-14Servicing Hydraulic Brake SystemReservoir ..................................................2-93Servicing Locations ...................................F2-29Servicing Oil System ................................2-92Servicing Oxygen System .........................2-98Servicing the Electric Toilet .......................2-95Single-Engine Before Landing ...................9-22Single-Engine Descent/Arrival ...................9-21Single Engine Go-Around .........................9-24Single Engine Landing Check ...................9-23Single Phase AC Electrical System ...........F2-23Spins ........................................................8-50Stalls ........................................................8-49Stall Speed ...............................................F8-2Stall Warning System ...............................2-56Standard, Alternate and Emergency Fuel . ...........T2-9Standby Magnetic Compass .....................2-84Starter Limitations ....................................5-8Subpanels ................................................F2-6Sun Visors ................................................2-65Surface Deicer System .............................2-52Symbols Definition ....................................8-5System Daily Preflight/Re-Arm Test ...................4-4


T

TACAN Control Panel andRange Indicator ................................... F3-17TACAN Control Panel andRange Indicator ................................... F3-17TACAN Systems .................................. 3-26Takeoff ................................................ 8-34Takeoff Distance (Flaps 0%) ................ 7-7Takeoff Weight .................................... 7-4Takeoff Weight To Achieve PositiveOne-Engine-Inop Climb ....................... 7-5Taxiing ................................................ 8-30Temperature Limits ............................. 5-30Three Phase AC Electrical System ......................F2-24Torque Limiter ..................................... 2-28Transponder Set (AN/APX-100) ........... 3-30, F3-25Trim Tabs ............................................ 2-41Turbulence and ThunderstormOperation ............................................ 8-56Turn-And-Slip Indicator ........................ 2-77, F3-13

U

UHF Command Set (AN/ARC-164) ....................3-8, F3-3Unpressurized Ventilation .................... 2-69Use of fuels ......................................... 2-91Use of Words Shall, Will, Should,and May .............................................. 1-12

V

Various Values for UTMGrid Coefficients .................................. T3-1Vertical Velocity Indicators ................... 2-81VHF AM Communications(VHF-20B) ........................................... 3-10, F3-4VHF AM-FM Command Set(AN/ARC-186) ..................................... 3-10, F3-11Voice Security System (TSEC/KY-28) .................3-12, F3-6Voice Security System (TSEC/KY-58) .................3-13VOR/LOC Navigation System .............. 3-23

W

Warning Annunciator Panel Legend.....................T2-6Warnings, Cautions, and Notes ........... 1-2Weather Radar Control-Indicator ......... F3-24

INDEX-5

TM 55-1510-219-10


W

Weather Radar Set (AN/APN-215) ............3-29Weight and Balance Clearance Form F...............6-8Weight Limitations ....................................5-19Wind Limitations .......................................5-32Windows ..................................................2-11Windshield Electrothermal Anti-Ice System . .........2-59Windshield Wipers ....................................2-62


W

Wind Swell Ditch Heading Evaluation ............ F9-4Wing Flap Extension Speeds ............... 5-24Wing Flaps .......................................... 2-42

Z

Zero Fuel Weight Limitation ................. 7-11

INDEX-6


CARL E. VUONOGeneral, United States Army

Official: Chief of Staff

PATRICIA P. HICKERSONColonel, United States Army

The Adjutant General

DISTRIBUTIUON:To Be distributed in accordance with DA Form 12-31-E, block no. 0121, -10 & CL maintenance

requirements for TM 55-1510-219-10.

*U.S. GOVERNMENT PRING OFFICE: 1994-300-421/02023

ELECTRONIC DA FORM 2028 INSTRUCTIONS

The following format must be used if submitting an electronic 2028. The subjectline must be exactly the same and all fields must be included; however, only thefollowing fields are mandatory: 1, 3, 4, 5, 6, 7, 8, 9, 10, 13, 15, 16, 17, and 27.From: ‘Whomever” <[email protected]>To: Is-lp @ redstone.army.milSubject: DA Form 20281. From: Joe Smith2. Unit: home3. Address: 4300 Park4. City: Hometown5. St: AL6. Zip: 777777. Date Sent: 19-OCT-938. Pub No: 55-2840-229-239. Pub Title: TM10. Publication Date: 04-JUL-8511. Change Number: 712. Submitter Rank: MSG13. Submitter Fname: Joe14. Submitter Mname: T15. Submitter Lname: Smith16. Submitter Phone: 123-123-123417. Problem: 118. Page: 219. Paragraph: 320. Line: 421. NSN: 522. Reference: 623. Figure: 724. Table: 825. Item: 928. Total: 12327. Text:This is the text for the problem below line 27.

PIN: 057981-000

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