HEADQUARTERS, DEPARTMENT · URGENT CHANGE NO. 5 TM 55-1510-221-10 C5 HEADQUARTERS DEPARTMENT OF THE...

TM 55-1510-221-10

TECHNICAL MANUAL WARNING DATA

OPERATOR’S MANUALFOR

ARMY RC-12H AIRCRAFTTABLE OF CONTENTS

INTRODUCTION

DESCRIPTION ANDOPERATION

AVIONICS

MISSION EQUIPMENT

OPERATING LIMITS ANDRESTRICTIONS

WEIGHT/BALANCE ANDLOADING

PERFORMANCE DATA

NORMAL PROCEDURES

EMERGENCY PROCEDURES“Approved for public release; distribution is unlimited.”

REFERENCES

HEADQUARTERS, DEPARTMENTOF THE ARMY

ABBREVIATIONS AND TERMS

30 DECEMBER 1988 ALPHABETICAL INDEX

This copy is a reprint which includes currentpages from Changes 1 through 4.

URGENT

CHANGE

NO. 5

TM 55-1510-221-10C5

HEADQUARTERSDEPARTMENT OF THE ARMY

WASHINGTON, D.C., 28 May 1998

OPERATOR’S MANUALFOR

ARMY RC-12H AIRCRAFT

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

TM 55-1510-221-10, 30 December 1988, 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 ii5-9 and 5-10510.1/(510.2 blank)8-28.1/(8-28.2 blank)8-29 and 8-30INDEX-3 and INDEX4

i and ii5-9 and 5-105-10.1/5-10.2 blank)8-28.1 and 8-28.28-29/(8-30 blank)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

04640

DENNIS J. REIMERGeneral, United States Army

Chief of Staff

DISTRIBUTION:To be distributed in accordance with Initial Distribution Number (IDN) 310993, requirements for TM 55

1510-221-10.

TM 55-1510-221-10C 4

CHANGE

NO. 4


WASHINGTON, D.C., 30 April 1993

Operator's ManualFor

Army RC-12H Aircraft

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


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


2-31 and 2-322-69 and 2-702-79 and 2-803-7 and 3-8

3-47 and 3-485-1 and 5-25-5 and 5-65-9 and 5-106-9 and 6-107-17 through 7-208-1 and 8-2

8-3 and 8-48-7 through 8-108-13 through 8-168-21 and 8-229-3 and 9-4

2-31 and 2-322-69 and 2-702-79 and 2-803-7 and 3-83-8.1 through 3-8.3/(3-8.4blank)

3-47 and 3-485-1 and 5-25-5 and 5-65-9 and 5-106-9 and 6-107-17 through 7-208-1 and 8-28-2.1/(8-2.2 blank)8-3 and 8-48-7 through 8-108-13 through 8-168-21 and 8-229-3 and 9-4


TM 55-1510-221-10

C 4


Official:

GORDON R. SULLIVANGeneral, United States Army

Chief of Staff

MILTON H. HAMILTONAdministrative Assistant to the

Secretary of the ArmyC4224

DISTRIBUTION:To be distributed in accordance with DA Form 12-31-E, block no 0993, require-

ments for TM 55-1510-221-10.

URGENTTM 55-1510-221-10

C 3

CHANGE

NO. 3


WASHINGTON, D.C., 17 August 1992




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


5-9 and 5-10- - - -8-17 through 8-20- - - -8-29 and 8-30

5-9 and 5-105-10.1/5-10.28-17 through 8-208-28.1/8-28.28-29 and 8-30



Official


Chief of Staff


Secretary of the Army02244

DISTRIBUTION:To be distributed in accordance with DA Form 12-31-E, block no. 0993, requirements

for TM 55-1510-221-10.

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

URGENT

TM 55-1510-221-10C 2

CHANGE

NO. 2


WASHINGTON, D.C., 30 June 1992




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


- - - - - 3-76A through 3-76F



Official:


Chief of Staff


Secretary of the Army01862

DISTRIBUTION:To be distributed in accordance with DA Form 12-31-E, block no. 0993, -10 & CL

maintenance requirements for TM 55-1510-221-10.

CHANGE

No. 1

TM 55-1510-221-10C 1


WASHINGTON, D.C., 9 November 1990




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


9-11 and 9-12 9-11 and 9-12- - - - 9-12.1/9-12.2



Official:

CARL E. VUONOGeneral, United States Army

Chief of Staff

THOMAS F. SIKORABrigadier General, United States Army

The Adjutant General

DISTRIBUTION:To be distributed in accordance with DA Form 12-31, -10 & CL Maintenance require-

ments for RC-12D Airplane, Reconnaissance.

TM 55-1510-221-10

WARNING PAGE

Personnel performing operations, procedures and practices which are included or implied in this technicalmanual shall observe the following warnings. Disregard of these warnings and precautionary information cancause injury or death.

NOISE LEVELS

Sound pressure levels in this aircraft during some operating conditions exceed the Surgeon General’s hear-ing conservation criteria, as defined in TM MED 501. Hearing protection devices, such as the aviator helmetor ear plugs shall 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 todo so may 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 nontoxic, it must be considered to be about the same as other freons and carbon dioxide, causing danger to per-sonnel primarily by reduction of oxygen available for proper breathing. During operation of the fire extin-guisher, ventilate personnel areas with fresh air. The liquid shall not be allowed to come into contact with theskin, as it may cause frostbite or low temperature burns because 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 of reflections of the light on the clouds.

CARBON MONOXIDE

When smoke, suspected carbon monoxide fumes, or symptoms of lack of oxygen (hypoxia) exist, all per-sonnel shall immediately 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 theskin. Do not allow them to remain on skin.

SERVICING AIRCRAFT

When conditions permit, the aircraft shall be positioned so that the wind will carry fuel vapors away fromall possible sources of ignition. The fueling unit shall maintain a distance of 20 feet between unit and fillerpoint. A minimum of 10 feet shall be maintained between fueling unit and aircraft.

a

TM 55-1510-221-10

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

SERVICING BATTERY

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

Corrosive Battery Electrolyte (Potassium Hydroxide). Wear rubber gloves, apron, and face shield whenhandling batteries. If potassium hydroxide is spilled on clothing, or other material wash immediately withclean water. If spilled on personnel, immediately start flushing the affected area with clean water. Continuewashing until medical assistance arrives.

JET BLAST

Occasionally, during starting, excess fuel accumulation in the combustion chamber causes flames to beblown from 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 itemspresent no radiation hazard to personnel unless seal has been broken due to aging or has accidentally been bro-ken. If seal is suspected 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 whenusing maintenance 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.

AUTOPILOT COMPATIBILITY

The RC-12H aircraft is certified with wingtip pods installed. Should the pods be removed, the autopilotsystem must be replaced with a standard C-12D autopilot. Effected wiring must also be changed.

b

TM 55-1516-221-10

TECHNICAL MANUAL HEADQUARTERSDEPARTMENT OF THE ARMY

WASHINGTON, D.C. 30 DECEMBER 1988

Operator’s ManualARMY MODEL RC-12H

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.

CHAPTER 1

CHAPTER 2

Section

CHAPTER 3.Section

CHAPTER 4.

Section

CHAPTER 5.

Section

CHAPTER 6. WEIGHT/BALANCE AND LOADING 6-1

I.II.

III.IV.V.

VI.VII.VIII.

IX.X.

XI.XII.

I.II.

Ill.IV.

1.II.

I.II.

Ill.IV.V.VI.

VII.VIII.

TABLE OF CONTENTS

INTRODUCTION

AIRCRAFT AND SYSTEMS DESCRIPTION ANDOPERATION

AircraftEmergency equipmentEngine and related systemsFuel systemsFlight controlsPropellersUtility systemsHeating, ventilation, cooling, and environmental controlsystemElectrical power supply and distribution systemLightingFlight instrumentsServicing, parking, and mooring

AVIONICSGeneralCommunicationsNavigationTransponder and radar

MISSION EQUIPMENT

Mission avionicsAircraft survivability equipment

OPERATING LIMITS AND RESTRICTIONS

GeneralSystem limitsPower limitsLoading limitsAirspeed limits, maximum and minimumManeuvering limitsEnvironmental restrictionsOther limitations

PAGE

1-1

2-1

2-12-192-202-272-342-422-512-52

2-582-672-702-77

3-13-13-2

3-203-77

4-1

4-14-1

5-1

5-15-15-65-85-6

5-105-105-11

Change 5 i

TM 55-1510-221-10

Section

TABLE OF CONTENTS (CONT’D)

1. GeneralII. Weight and balanceIII. Fuel/oilIV. Cargo loadingV. Center of gravity

CHAPTER 7. PERFORMANCE DATA

CHAPTER 8. NORMAL PROCEDURES

Section I. Mission planningII. Operating procedures and maneuvers

Ill. Instrument flightIV. Flight characteristicsV. Adverse environmental conditionsVI. Crew duties

CHAPTER 9. EMERGENCY PROCEDURES

Section I. Aircraft systems

APPENDIX A. REFERENCES

APPENDIX B. ABBREVIATIONS AND TERMS

INDEX

PAGE

6-16-1

6-146-146-14

7-1

8-1

8-18-1

8-238-248-268-29

9-1

9-1

A-1

B-1

INDEX-1

ii Change 5

TM 55-1510-221-10

CHAPTER 1

INTRODUCTION

1-1. GENERAL. 1-5. APPENDIX B, ABBREVIATIONS ANDTERMS.

These instructions are for use by the operator(s).They apply to the RC-12H aircraft. Appendix B is a listing of abbreviations and

terms used throughout the manual.

1-2. WARNINGS, CAUTIONS, AND NOTES.

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

1-6. INDEX.

The index lists, in alphabetical order, everytitled paragraph, figure, and table contained in thismanual. Chapter 7, Performance Data, has an addi-tional index within the chapter.

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

1-7. ARMY AVIATION SAFETY PROGRAM.

Reports necessary to comply with the safety pro-gram are prescribed in AR 385-40.

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

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

For information concerning destruction of Armymateriel to prevent enemy use, refer to TM 750-244-1-5.

NOTE 1-9. FORMS AND RECORDS.

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

Army aviators flight record and aircraft mainte-nance records which are to be used by crew mem-bers are prescribed in DA PAM 738-751 and TM55-1600-342-23.

1-3. DESCRIPTION.

This manual contains the best operating instruc-tions and procedures for the RC-12H aircraft undermost circumstances. The observance of limitations,performance, and weight/balance data provided ismandatory. The observance of procedures is manda-tory except when modification is required becauseof multiple emergencies, adverse weather, terrain,etc. Your flying experience is recognized, and there-fore, basic flight principles are not included. THISMANUAL SHALL BE CARRIED IN THE AIR-CRAFT AT ALL TIMES.

1-10. EXPLANATION OF CHANGE SYMBOLS.

1-4. APPENDIX A, REFERENCES.

Appendix A is a listing of official publicationscited within the manual applicable to and availablefor flight crews.

Changes, except as noted below, to the text andtables, including new material on added pages, areindicated by a vertical line in the outer marginextending close to the entire area of the materialaffected; exception: pages with emergency markings,which consist of black diagonal lines around threeedges, may have the vertical line or change symbolplaced along the inner margins. Symbols show cur-rent changes only. A miniature pointing hand sym-bol is used to denote a change to an illustration.However, a vertical line in the outer margin, ratherthan miniature pointing hands, is utilized whenthere have been extensive changes made to an illus-tration. Change symbols are not utilized to indicatechanges in the following:

1-1

TM 55-1510-221-10

a. Introductory material.

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

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

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

1-11. AIRCRAFT DESIGNATION SYSTEM.

The designation system prescribed by AR 70-50is used in aircraft designations as follows:

EXAMPLE RC-12H

Within this technical manual the word “shall” isused to indicate a mandatory requirement. Theword “should” is used to indicate a nonmandatorybut preferred method of accomplishment. The word“may” is used to indicate an acceptable method ofaccomplishment. The word “will” is used to expressa declaration of purpose and may also be used wheresimple futurity is required.

1-13. PLACARD ITEMS.

R - Modified mission symbol (Reconnais- All placard items (switches, controls, etc.) aresance) shown throughout this manual in capital letters.

C - Basic mission and type symbol (cargo)

12 - Design number

H - Series symbol

1-12. USE OF WORDS SHALL, WILL, SHOULD,AND MAY.

1-2

TM 55-1510-221-10

CHAPTER 2

AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATION

2-1. INTRODUCTION.

The purpose of this chapter is to describe theaircraft and its systems and controls which contrib-ute to the physical act of operating the aircraft. Itdoes not contain descriptions of avionics or missionequipment, covered elsewhere in this manual. Thischapter contains descriptive information and doesnot describe procedures for operation of the aircraft.These procedures are contained within appropriatechapters in the manual. This chapter also containsthe emergency equipment installed. This chapter isnot designed to provide instructions on the completemechanical and electrical workings of the varioussystems; therefore, each is described only in enoughdetail to make comprehension of that system suffi-ciently complete to allow for its safe and efficientoperation.

2-2. GENERAL.

The RC-12H is a pressurized, low wing, allmetal aircraft, powered by two PT6A-41 turbopropengines (fig. 2-1 and 2-11), and has all weather capa-bility. Distinguishable features of the aircraft are theslender, streamlined engine nacelles, an aft rotatingboom antenna, mission antennas, wing tip pods, aT-tail and a ventral tin below the empennage. Thebasic mission of the aircraft is radio reconnaissance.Cabin entrance is made through a stair-type door(fig. 2-2) on the left side of the fuselage.

2-3. DIMENSIONS.

Overall aircraft dimensions are shown in figure2-3.

2-4. GROUND TURNING RADIUS.

Minimum ground turning radius of the aircraftis shown in figure 2-4.

2-5. MAXIMUM WEIGHTS.

Maximum takeoff gross weight is 15,000pounds. Maximum landing weight is 15,000 pounds.Maximum ramp weight is 15,090 pounds. Maxi-mum zero fuel weight is 11,500 pounds.

AIRCRAFTSection I.

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 maintainedwith engines operating at idle are also shown. Tem-perature and velocity of exhaust gases at varyinglocations aft of the exhaust stacks are shown formaximum power. The danger area extends to 40 feetaft of the exhaust stack outlets. Propeller dangerareas are also shown.

2-7. LANDING GEAR SYSTEM.

The landing gear is a retractable, tricycle type,electrically operated by a single DC motor. Thismotor drives the main landing gear actuatorsthrough a gear box and torque tube arrangement,and also drives a chain mechanism which controlsthe position of the nose gear. Positive down-locksare installed to hold the drag brace in the extendedand locked position. The down-locks are actuated byovertravel of the linear jackscrews and are held inposition by a spring-loaded overcenter mechanism.The jackscrew in each actuator holds all three gearsin the UP position, when the gear is retracted. Afriction clutch between the gearbox and the torqueshafts protects the motor from electrical overload inthe event of a mechanical malfunction. A 150-ampere current limiter, located on the DC distribu-tion bus under the center floorboard, protectsagainst electrical overload. Gear doors are openedand closed through a mechanical linkage connectedto the landing gear. The nose wheel steering mecha-nism is automatically centered and the rudder ped-als relieved of the steering load when the landinggear is retracted. Air-oil type shock struts, filled withcompressed air and hydraulic fluid, are incorporatedwith the landing gear. Gear retraction or extensiontime is approximately six seconds.

(1.) Landing Gear Control Switch. Land-ing gear system operation is controlled by a manu-ally actuated, wheel-shaped switch placarded LDGGEAR CONTR - UP - DN, located on the left sub-panel (fig. 2-6). The control switch and associatedrelay circuits are protected by a S-ampere circuitbreaker, placarded LANDING GEAR RELAY onthe overhead circuit breaker panel (fig. 2-26).

2-1

TM 55-1510-221-10

1. Weather radar2. Air conditioner condenser air outlet3. Nose avionics compartment access door4. Windshield wipers5. AN/APR39 spiral antenna6. VHF/AM/FM antenna7. Global positioning system antenna8. Low band vert bent blade (upper) antenna9. Transponder antenna10. VHF comm antenna11. “P” band antenna12. Static air source13. Relief tube drain14. ELT control switch15. UHF L-band antenna16. Emergency light17. Cargo door18. Cabin door19. Strobe light20. Low band vert bent blade (lower) antenna21. Navigation light22. Ice light23. Outside air temperature gage probe24. Wide band data link fwd antenna25. Glideslope antenna AP 011758

Figure 2-1. General Exterior Arrangement (Sheet 1 of 5)

2-2

TM 55-1510-221-10

26. Navigation light27. Strobe beacon28. Dorsal fin (ADF sense) antenna29. AN/APR-44 antenna30. Flare dispenser31. Mid-band dipole antenna32. Emergency entrance/exit hatch33. TACAN antenna34. Nose avionics compartment access door35. Air conditioner condenser air inlet36. Radome37. High band vert & horiz antenna38. Ice light39. AN/APR-39 blade antenna40. High band monopole antenna41. Static air source42. Oxygen system servicing door43. “P” band antenna44. Wide band data link aft antenna45. Mid-band dipole antenna46. Static wick47. Aft rotating boom antenna


AP 011758.1)

2-3

TM 55-1510-221-10

1. AN/APR-39 spiral antenna2. UHF comm & intercept antenna3. Mid-band dipole antenna4. Tailet5. Flare dispenser6. Low band horiz towel bar antenna7. AN/APR-44 antenna8. VOR NAV/LOC antenna9. Strobe antenna10. Navigation light11. ELINT & DF antenna pod12. Recognition light13. Stall warning vane14. High band monopole antenna15. Bleed air heat exchanger air inlet16. Pitot tube17. Wide band data link fwd antenna18. Taxi light19. Landing lights

AP 011758.2


2-4

TM 55-1510-221-10

1. UHF comm & intercept antenna2. AN/APR-39 spiral antenna3. Navigation light4. ELINT & DF antenna pod5. Fuel filler cap6. Mid-band dipole antenna7. AN/APR-39 antenna8. Marker beacon antenna9. Wide band data link antenna10. High band vert & horiz antenna11. TACAN antenna12. GPS antenna13. Low band vert bent blade antenna (upper)14. High band monopole antenna15. Strobe beacon16. Low band horiz towel bar antenna17. UHF L-band antenna18. VHF comm antenna19. Transponder antenna AP 011758.3


2-5

TM 55-1510-221-10

1. Radio altimeter antenna2. VHF/AM/FM antenna3. Glideslope antenna4. ADF loop antenna5. Fuel filler cap6. Mid-band dipole antenna7. AN/APR-39 spiral antenna8. ELINT & DF antenna pod9. Navigation light10. UHF comm & Intercept antenna11. High band monopole antenna12. INS TACAN antenna13. AN/APR-44 antenna14. ELT antenna15. Static wick16. Strobe beacon17. Dorsal fin (ADF sense antenna)18. Strobe dams AP 011758.4


2-6

TM 55-1510-221-10

Figure 2-2. General Interior Arrangement

2-7

TM 55-1510-221-10

Figure 2-3. Principal Dimensions

2-8

Figure 2-4. Ground Turning Radius

2-9

TM 55-1510-221-10

Figure 2-5. Exhaust and Propeller Danger Areas

2-10

TM 55-1510-221-10

CAUTION/ADVISORY-ANNUNCIATOR PANEL

Figure 2-6. Subpanels

2-11

TM 55-1510-221-10

a. Landing Gear Down Position-IndicatorLights. Landing gear down position is indicated bythree green lights on the left subpanel, placardedGEAR DOWN. These lights may be checked byoperating the ANNUNCIATOR TEST switch. Thecircuit is protected by a 5-ampere circuit breaker,placarded LANDING GEAR IND, on the overheadcircuit breaker panel (fig. 2-26).

b. Landing Gear Position Warning Lights.Two red bulbs, wired in parallel and activated bymicroswitches independent of the GEAR DOWNposition indicator lights, are positioned inside theclear plastic grip on the landing gear control switch.These lights illuminate whenever the landing gearswitch is in either the UP or DN position and thegear is in transit. Both bulbs will also illuminateshould either or both power levers be retarded belowapproximately 79 to 81% N1 when the landing gearis not down and locked. To turn the switch lightsOFF during single-engine operation, the power leverfor the inoperative engine must be advanced to aposition which is higher than the setting of the warn-ing horn microswitch. Extending the landing gearwill also turn the lights off. Both red lights indicatethe same warning conditions, but two are providedfor a fail-safe indication in the event one bulb burnsout. The circuit is protected by a 5-ampere circuitbreaker, placarded LANDING GEAR IND, on theoverhead circuit breaker panel (fig. 2-26).

c. Landing Gear Warning Light Test Button.A test button, placarded HDL LT TEST, is locatedon the left subpanel. Failure of the landing gearswitch to illuminate red, when this test button ispressed, indicates two defective bulbs or a circuitfault. The circuit is protected by a 5-ampere circuitbreaker, placarded LANDING GEAR RELAYCONTROL, on the overhead circuit breaker panel(fig. 2-26).

d. Landing Gear Warning Horn. When eitherpower lever is retarded below approximately 79 to81 N, when the landing gear is not down and lockedor if the flaps are extended beyond 40% and thelanding gear is not down and locked, a warning hornlocated in the overhead control panel will soundintermittently. To prevent the warning horn fromsounding during long descents or an ILS approach,a pressure differential “Q” switch is connected intothe copilot’s static line. The switch prevents thewarning horn from sounding until airspeed dropsbelow 140 KIAS. An altitude sensing switch isinstalled in series with the 140 KIAS “Q“ switch

which prevents the warning horn from soundingafter climbing through 12,500 feet MSL. The hornwill be engaged when the aircraft descends through10,500 MSL. The warning horn circuit is protectedby a 5-ampere circuit breaker, placarded LANDINGGEAR WARN, on the overhead circuit breakerpanel (fig. 2-26).

e. Landing Gear Warning Horn Test Switch.The landing gear warning horn may be tested by thetest switch on the right subpanel (fig. 2-6). Theswitch, placarded STALL WARN TEST - OFF -LDG GEAR WARN TEST, will sound the landinggear warning horn and illuminate the landing gearposition warning lights when moved to the momen-tary LDG GEAR WARN TEST position. The circuitis protected by a 5-ampere circuit breaker, placardedLANDING GEAR WARN, on the overhead circuitbreaker panel (fig. 2-26).

f. Landing Gear Safety Switches. A safetyswitch on each main landing gear shock strut con-trols the operation of various aircraft systems thatfunction only during flight or only during groundoperation. These switches are mechanically actuatedwhenever the main landing gear shock struts areextended (normally after takeoff), or compressed(normally after landing). The safety switch on theright main landing gear strut activates the landinggear control circuits, cabin pressurization circuitsand the flight hour meter flight time function on thecopilot’s clock when the strut is extended. Thisswitch also activates a down-lock hook, preventingthe landing gear from being raised while the aircraftis on the ground. The hook, which unlocks automat-ically after takeoff, can be manually overridden bypressing down on the red button, placarded DNLOCK REL located adjacent to the landing gearswitch. If the override is used and the landing gearcontrol switch is raised, power will be supplied tothe warning horn circuit and the horn will sound.The safety switch on the left main landing gear strutactivates the left and right engine ambient air shut-off valves when the strut is extended.

g. Landing Gear Alternate Engage Handle.During manual landing gear extension, the landinggear motor must be disengaged from the landinggear drive mechanism. This is accomplished by amanually-operated clutch disengage lever (fig. 2-7)located adjacent to the landing gear alternate exten-sion handle (fig. 2-7). To disengage the clutch, pullthe alternate engage handle up and turn clockwise.To engage the clutch, turn the alternate engage han-dle counterclockwise and release.

2-12

TM 55-1510-221-10

1. Power levers2. Propeller levers3. Condition levers4. FLAP cover5. Rudder trim6. Flare dispenser control panel7. Flare dispenser switch8. TACAN Control panel9. NAV 1/NAV 2 control panel10. Elev trim & rudder boost switches11. Marker beacon audio control panel12. Mode selector unit13. ADF control panel14. VHF AM/FM control panel15. HF command set control panel (KHF-950)16. No. 2 UHF command set17. Transponder control panel18. VHF AM control panel19. No. 1 UHF command set20. Flight director mode select panel (MS 500)21. Autopilot pitch and turn control22. No. 1 HSI course and heading knobs23. Emergency Ianding gear alternateengage handle24. Emergency landing gear extensionratchet handle25. Aileron tab control and position indicator26. Elevator tab control and position Indicator27. Go-around button

AP 011769

Figure 2-7. Control Pedestal

2-13

TM 55-1510-221-10

Continued pumping of handle afterGEAR DOWN position indicator lights(3) are illuminated could damage thedrive mechanism, and prevent subsequentgear retraction.

h. Landing Gear Alternate Extension Handle.Manual landing gear extension is provided througha manually powered system as a backup to the elec-trically operated system. Before manually extendingthe gear, make certain that the landing gear switchis in the down position with the LANDING GEARRELAY circuit breaker pulled. Pulling up on thealternate engage handle, located on the floor, andturning it clockwise will lock it in that position.When the alternate engage handle is pulled, themotor is electrically disconnected from the systemand the alternate drive system is locked to the gear-box and motor. When the alternate drive is lockedin, the chain is driven by a continuous action ratchetwhich is activated by pumping the landing gearalternate extension handle adjacent to the alternateengage handle.

k. Wheel Brake System. The main wheels areequipped with multiple-disc hydraulic brakes actu-ated by master cylinders attached to the rudder ped-als at the pilot’s and copilot’s position. Braking ispermitted from either set of rudder pedals. Brakefluid is supplied to the system from the reservoir inthe nose compartment. The toe brake sections of therudder pedals are connected to the master cylinderswhich actuate the system for the correspondingwheels. No emergency brake system is provided.Repeated and excessive application of brakes, with-out allowing sufficient time for cooling to accumu-late between applications, will cause a loss of brak-ing efficiency, possible failure of brake or wheelstructure, possible blowout of tires, and in extremecases may cause the wheel and brake assembly to bedestroyed by fire.

(1.) After a manual landing gear extensionhas been made, do not stow the handle, move anylanding gear controls, or reset any switches or circuitbreakers. The gear cannot be retracted manually.

2-8. PARKING BRAKE.

(2.) After a practice manual extension, thealternate handle may be stowed and the landing gearretracted electrically. Rotate the alternate engagehandle counterclockwise and push it down. Stow thehandle, push in the LANDING GEAR RELAY cir-cuit breaker on the overhead circuit breaker paneland retract the gear in the normal manner with thelanding gear switch. Refer to Chapter 9 for emer-gency gear extension procedures.

Dual parking brake valves are installed belowthe cockpit floor. Both valves can be closed simulta-neously by pressing both brake pedals to build uppressure, then pulling out the handle placardedPARKING BRAKE, on the left subpanel. Pullingthe handle full out sets the check valves in the sys-tem and any pressure being applied by the toebrakes is maintained. Parking brakes are releasedwhen the brake handle is pushed in. The parkingbrake may be set from either cockpit position. Park-ing brakes shall not be set during flight.

i. Tires. The aircraft is equipped with dual 22x 6.75 - 10 tubeless, 8-ply rated, rim-inflation tireson each main gear. The nose gear is equipped witha single 22 x 6.75 - 10, 8-ply rated, tubeless, rim-inflated tire.

2-9. ENTRANCE AND EXIT PROVISIONS.

NOTE

j. Steerable Nose Wheel. The aircraft can be Two keys are provided in the loose toolsmaneuvered on the ground by the steerable nose and equipment bag. Both keys will fit thewheel system. Direct linkage from the rudder pedals locks on the cabin door, emergency hatch,(fig. 2-8) to the nose wheel steering linkage allows tailcone access door and the right and leftthe nose wheel to be turned 12° to the left of center nose avionics compartment doors.

or 14° to the right. When rudder pedal steering isaugmented by the main wheel braking action, thenose wheel can be deflected up to 48° either side ofcenter. Shock loads which would normally be trans-mitted to the rudder pedals are absorbed by a springmechanism in the steering linkage. Retraction of thelanding gear automatically centers the nose wheeland disengages the steering linkage from the rudderpedals.

2-14

TM 55-1510-221-10

1. Free air temperature gage2. Oxygen system pressure gages3. Storm window lock4. Oxygen regulator control panel5. Control wheel6. Sun visor7. Overhead circuit breaker and control panel

Figure 2-8. Cockpit

8. Windshield wiper9. Magnetic compass10. Rudder pedals11. Mission control panel12. Pedestal extension13. Assist step

AP 012087

2-15

TM 55-1510-221-10

a. Cabin Door.

Structural damage may be caused if morethan one person is on the cabin door at atime. The door is weight limited to 300pounds or less.

A swing-down door (fig. 2-9), hinged at the bot-tom, provides a stairway for normal and emergencyentrance and exit. Two of the steps are movable andfold flat against the door in the closed position. Astep folds down over the door sill when the dooropens to provide a platform (step) for door seal pro-tection. A plastic encased cable provides support forthe door in the open position, a handhold, and aconvenience for closing the door from inside. Ahydraulic damper permits the door to lower gradu-ally during opening. A rubber seal around the doorseals the pressure vessel while the aircraft is inflight. The door locking mechanism is operated byeither of the two mechanically interconnected han-dles, one inside and the other outside the door.When either handle is rotated, three rotating cam-type latches on either side of the door capture postsmounted on the cargo door. In the closed position,the door becomes an integral part of the cargo door.A button adjacent to the door handle must bedepressed before the handle can be rotated to openthe door. A bellows behind the button is inflatedwhen the aircraft is pressurized to prevent acciden-tal unlatching and/or opening of the door. A smallround window just above the second step permitsobservation of the pressurization safety bellows. Aplacard adjacent to the window instructs the opera-tor that the safety lock arm is in position around thebellows shaft which indicates a properly lockeddoor. Pushing the red button adjacent to the windowwill illuminate the inside door mechanism. ACABIN DOOR annunciator light on the caution/advisory panel will illuminate if the door is notclosed and all latches fully locked. The cabin dooropening is 21.5 inches wide by 50.0 inches high.

b. Cargo Door. A swing-up door (fig. 2-9)hinged at the top, provides cabin access for loadingcargo or bulky items. After initial opening force isapplied, gas springs will completely open the cargodoor automatically. The door is counterbalancedand will remain in the open position. A door sup-port rod is used to hold the door in the open posi-tion, and to aid in overcoming the pressure of thegas spring assemblies when closing the door. Onceclosed, the gas springs apply a closing force to assistin latching the door. A rubber seal around the doorseals the pressure vessel while in flight. The door

2-16

locking mechanism is operated only from inside theaircraft, and is operated by two handles, one in thebottom forward portion of the door and the other inthe upper aft portion of the door. When the upperaft handle is operated per placard instructions, tworotating 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.The bottom handle, when operated per placardinstructions, actuates four pin lug latches across thebottom of the door. A button on the upper aft han-dle must be pressed before the handle can bereleased to open or latch the door. A latching leveron the bottom handle must be lifted to release thehandle before the lower latches can be opened.These act as additional aids in preventing accidentalopening or unlatching of the door. The cabin andcargo doors are equipped with dual sensing circuitsto provide the crew remote indication of cabin/cargodoor security. An annunciator light placardedCABIN DOOR will illuminate if the cabin or cargodoor is open and the BATT switch in ON. If the bat-tery switch is OFF, the annunciator will illuminateonly if the cargo door is not securely closed andlatched. The cargo door sensing circuit receivespower from the hot battery bus. The cargo dooropening is 52.0 inches wide by 52.0 inches high.

Insure that the cabin door is closed andlocked. Operating the cargo door whilethe cabin door is open may damage thedoor hinge and adjacent structure.

(1.) Opening cargo door.

Avoid side loading of the gas springs toprevent damage to the mechanism.

1.

2.

3.

4.

5.

Handle access door (lower for-ward corner of door) -Unfastenand open.

Handle - Lift hook and move toOPEN position.

Handle access door - Secure.

Handle access door (upper aft cor-ner of door) - Unfasten and open.

Handle - Press button and lift toOPEN position then latch inplace.

TM 55-1510-221-10

Figure 2-9. Cabin and Cargo Doors

2-17

TM 55-1510-221-10

6.

7.

8.

9.

10.


Door support rod - Attach oneend to cargo door ball stud (onforward side of door).

Support rod detent pin - Check inplace.

Cabin door sill step - Push out onand allow cargo door to swingopen. Gas springs will automati-cally open the door.

Door support rod - Attach freeend to ball stud on forward fuse-lage door frame.

(2.) Closing cargo door.

Avoid side loading of the gas springs toprevent damage to the mechanism.

1.

2.

3.

4.

5.

6.

7.

8.

9.

Door support rod - Detach fromfuselage door frame ball stud,then firmly grasp free end of rodwhile exerting downward force toovercome the pressure of gasspring assemblies. Then removesupport rod from door as gasspring assemblies pass over-centerposition.

Cargo door - Pull closed, usingfinger hold cavity in fixed cabindoor step.

Handle access door (upper aft cor-ner of door) - Unfasten and open.

Handle - Press button and pullhandle down until it latches inclosed position.


Handle access door (lower for-ward comer of door) -Unfastenand open.

Handle - Move to full forwardposition.

Safety hook - Check locked inposition by pulling aft on handle.


c. Cabin Emergency Hatch. The cabin emer-gency hatch, placarded EXIT - PULL, is located onthe right cabin sidewall just aft of the copilot’s seat.

2-16

The hatch may be released, from the inside with apull-down handle. A flush mounted pull out handleallows the hatch to be released from the outside. Thehatch is of the non-hinged plug type, which removescompletely from the frame when the latches arereleased. The hatch can be key locked from theinside, to prevent opening from the outside. Theinside handle will unlatch the hatch whether or notit is locked, by overriding the locking mechanism.The keylock should be unlocked prior to flight toallow removal of the hatch from the outside in theevent of an emergency. The key remains in the lockwhen the hatch is locked and can be removed onlywhen the hatch is unlocked. The key slot is in thevertical position when the hatch is unlocked.Removal of the key from the lock before flightassures the pilot that the hatch can be removed fromthe outside if necessary.

d. Cabin Door Caution Light. As a safety pre-caution, two illuminated MASTER CAUTIONlights, on the glare shield and a steadily illuminatedCABIN DOOR yellow caution annunciator light onthe caution/advisory panel indicate the cabin door isnot closed and locked. This circuit is protected by5-ampere circuit breakers placarded ANN PWR andANN IND, located on the overhead circuit breakerpanel (fig. 2-26).

2-10. WINDOWS.

a. Cockpit Window. The pilot and copilothave side windows, a windshield and storm win-dows, which provide visibility from the cockpit. Thestorm windows may be opened on the ground orduring unpressurized flight.

b. Cabin Windows. The outer cabin windowshave two-ply construction, are of the pressure typeand are integral parts of the pressure vessel. Allcabin windows are painted over except for the win-dow farthest aft on the right side and the windowfarthest aft on the left side. All unpainted windowshave flaps which may be raised to permit visibilityor lowered to black out the windows.

2-11. SEATS.

a. Pilot and Copilot Seats. The pilot and copi-lot seats (fig. 2-10) are separated from the cabin bymovable curtains. The controls for vertical heightadjustment and fore and aft travel are located undereach seat. The forward and aft adjustment handle islocated beneath the bottom front inboard corner ofeach seat. Pulling up on the handle allows the seatto move fore or aft. The height adjustment handle islocated beneath the bottom front outboard corner ofeach seat. Pulling up on the handle, allows the seat

TM 55-1510-221-10

1. Adjustable headrest2. Seatbelt/shoulder harness buckle3. Moveable armrest4. Seat height adjustment (pilot), foreand aft adjustment (copilot)5. Seat fore and aft adjustment (pilot),height adjustment (copilot)6. Expandable map pocket

AP 004766

Figure 2-10. Pilot and Copilot Seats

to move up and down. Both seats have adjustableheadrests and armrests which will raise and lowerfor access to the cockpit. Handholds on either sideof the overhead panels and a fold-away protectivepedestal step are provided for pilot and copilot entryinto the cockpit. For the storage of maps and theoperator’s manual, pilot and copilot seats have aninboard-slanted, expandable pocket affixed to thelower portion of the seat back. Pocket openings areheld closed by shock cord tension.

b. Pilot and Copilot Seat Belts and ShoulderHarnesses. Each pilot and copilot seat is equipped

with a lap-type seat belt and shoulder harness con-nected to an inertia reel. The shoulder harness beltis of the “Y” configuration with the single strapbeing contained in an inertia reel attached to thebase of the seatback. The two straps are worn withone strap over each shoulder and fastened by metalloops into the seat belt buckle. The spring loading atthe inertia reel keeps the harness snug but will allownormal movement required during flight operations.The inertia reel is designed with a locking devicethat will secure the harness in the event of suddenforward movement or an impact action.

Section II. EMERGENCY EQUIPMENT

2-12. DESCRIPTION.

The equipment covered in this section includesall emergency equipment, except that which formspart of a complete system. For example, landinggear system, etc. Chapter 9 describes the operationof emergency exits and location of all emergencyequipment.

2-13. FIRST AID KITS.

Four first aid kits are included in the survivalkit.

2-19

TM 55-1510-221-10

2-14. HAND-OPERATED FIRE EXTINGUISHER.

Repeated or prolonged exposure to highconcentrations of monobromotrifluo-romethane (CF 3Br) or decompositionproducts should be avoided. The liquidshall not be allowed to come into contactwith the skin, as it may cause frost bite orlow temperature burns because of its verylow boiling point.

One hand-operated fire extinguisher is mounted Tie-down provisions for a survival raft and kitbelow the pilot’s seat and a second extinguisher is are provided just forward of the toilet on the rightlocated on the left cabin sidewall, aft of the cabin hand side of the cabin (fig. 2-2).

door. They are of the monobromotrifluromethane(CF3Br) type. The extinguisher is charged to a pres-sure of 150 to 170 PSI and emits a forceful stream.Use an extinguisher with care within the limitedarea of the cabin to avoid severe splashing.

NOTE

Engine tire extinguisher systems aredescribed in Section III.

2-15. SURVIVAL KITS.

Section III. ENGINES AND RELATED SYSTEMS

2-16. ENGINES.

The aircraft is powered by two PT6A-41 turbo-prop engines (fig. 2-11). The engine has a three stageaxial, single stage centrifugal compressor, driven bya single stage reaction turbine. The power turbine, atwo stage reaction turbine, counter-rotating with thecompressor turbine, drives the output shaft. Boththe compressor turbine and the power turbine arelocated in the approximate center of the engine withtheir shafts extending in opposite directions. Beinga reverse flow engine, the ram air supply enters thelower portion of the nacelle and is drawn in throughthe aft protective screens. The air is then routed intothe compressor. After it is compressed, it is forcedinto the annular combustion chamber, and mixedwith fuel that is sprayed in through 14 nozzlesmounted around the gas generator case. A capaci-tance discharge ignition unit and two spark igniterplugs are used to start combustion. After combus-tion, the exhaust passes through the compressor tur-bine and two stages of power turbine then is routedthrough two exhaust ports near the front of theengine. A pneumatic fuel control system schedulesfuel flow to maintain the power set by the gas gener-ator power lever. The accessory drive at the aft endof the engine provides power to drive the fuelpumps, fuel control, the oil pumps, the refrigerantcompressor (right engine), the starter-generator, and

the turbine tachometer transmitter. The reductiongearbox forward of the power turbine provides gear-ing for the propeller and drives the propellertachometer transmitter, the propeller overspeed gov-ernor, and the propeller governor.

2-17. ENGINE COMPARTMENT COOLING.

The forward engine compartment including theaccessory section is cooled by air entering aroundthe exhaust stack cutouts, the gap between the pro-peller spinner and forward cowling, and exhaustingthrough ducts in the upper and lower aft cowling.

2-18. AIR INDUCTION SYSTEMS - GENERAL.

Each engine and oil cooler receives ram air duc-ted from an air scoop located within the lower sec-tion of the forward nacelle. Special components ofthe engine induction system protect the power plantfrom icing and foreign object damage.

2-19. FOREIGN OBJECT DAMAGE CONTROL.

The engine has an integral air inlet screendesigned to obstruct objects large enough to damagethe compressor.

2-20

TM 55-1510-221-10

1. Primary prop governor2. Torque pressure transmitter3. Torque pressure switch4. Torque pressure manifold5. Exhaust duct6. TGT temperature probe7. Fuel flow divider manifold8. Fire detector9. Engine mount bolt

10. Engine mount truss assembly11. Engine air intake screen12. Ignition exciter13. Starter-generator14. Fuel boost pump

15. Air conditioner compressor drivebelt (#2 engine only)16. Fire detector17. Air conditioner compressor(#2 engine only)18. Bleed air adapter19. Bleed air line20. Engine mount21. Ignition exciter plug22. Oil scavenge tubes23. Overspeed governor24. Prop deice brush block bracket25. Prop reverse linkage lever

AP 005485.1Figure 2-11. PT6A-41 Engine (Sheet 1 of 2)

2-21

TM 55-1510-221-10

26. Fuel control unit27. Fuel control unit control rod28. Starter generator leads29. Engine driven fuel pump30. Power control lever31. Prop interconnect linkage (aft)32. Oil pressure transducer33. Engine mount34. Fireshield35. Trim resistor thermocouple36. Prop interconnect linkage (fore)37. Prop shaft38. Tach generator

AP 005485.2

39. Chip detector40. Oil pressure tube41. Fire extinguisher line42. Ignition exciter plug43. Engine mount bolt44. Linear actuator45. Engine baffle and seal assy46. Fuel/oil heater47. Tach-generator (aft)48. Drain manifold49. Overhead breather tube60. Engine truss mounting bolt

Figure 2-11. PT6A-41 Engine (Sheet 2 of 2)

2-22

2-20. ENGINE ICE PROTECTION SYSTEMS.

a. Inertial Separator.

After the ice vanes have been manuallyextended, they may be mechanically actu-ated only. No electrical extension orretraction shall be attempted as damageto the actuator may result. Linkage in thenacelle area must be reset prior to opera-tion of the electric system.

An inertial separation system is built into eachengine air inlet to prevent moisture particles fromentering the engine inlet plenum under icing condi-tions. A movable vane and a bypass door are low-ered into the airstream when operating in visiblemoisture at 5°C or colder, by energizing electricalactuators with the switches, placarded ICE VANE -RETRACT - EXTEND, located on the overheadcontrol panel. A mechanical backup system is pro-vided, and is actuated by pulling the T-handles justbelow the pilot’s subpanel placarded ICE VANE -No. 1 ENG - No. 2 ENG. Decrease airspeed to 160knots or less to reduce forces for manual extension.Normal airspeed may then be resumed.

(1.) The vane deflects the ram airstreamslightly downward to introduce a sudden turn in theairstream to the engine, causing the moisture parti-cles to continue on undeflected, because of theirgreater momentum, and to be discharged overboard.

(2.) While in the icing flight mode, theextended position of the vane and bypass door isindicated by green annunciator lights, No. 1 VANEEXT and No. 2 VANE EXT.

(3.) In the non-ice protection mode, thevane and bypass door are retracted out of the air-stream by placing the ice vane switches in theRETRACT position. The green annunciator lightswill extinguish. To assure adequate oil cooling,retraction should be accomplished at 15°C andabove. The vanes should be either extended orretracted; there are no intermediate positions.

(4.) If for any reason the vane does notattain the selected position within 15 seconds, a yel-low No. 1 VANE FAIL or No. 2 VANE FAIL lightilluminates on the caution/advisory panel. In thisevent, the manual backup system should be used.When the vane is successfully positioned with themanual system, the yellow annunciator lights willextinguish. During manual system use, the electricmotor switch position must match the manual han-dle position for a correct annunciator readout.

TM 55-1510-221-10

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 plumbingto the inlet lip. The gas flows through the inside ofthe lip to the bottom where it is allowed to escape.

(2.) Fuel heater. An oil-to-fuel heatexchanger, located on the engine accessory case,operates continuously and automatically to heat thefuel sufficiently to prevent ice from collecting in thefuel control unit. Each fuel control unit is protectedagainst ice. Fuel control heat is automatically turnedon for all engine operations.

2-21. ENGINE FUEL CONTROL SYSTEM.

a. Description. The basic engine fuel systemconsists of an engine driven fuel pump, a fuel con-trol unit, a fuel flow divider, a dual fuel manifoldand fourteen fuel nozzles. The fuel flow divider actsas a drain valve to clear residual fuel after engineshutdown.

b. Fuel Control Unit. One fuel control unit ismounted on the accessory case of each engine. Thisunit is a hydro-pneumatic metering device whichdetermines the proper fuel schedule for the engine toproduce the amount of power requested by the rela-tive position of its power lever. The control of devel-oped engine power is accomplished by adjusting theengine compressor turbine (N1) speed. N1 speed iscontrolled by varying the amount of fuel injectedinto the combustion chamber through the fuel noz-zles. Engine shutdown is accomplished by movingthe appropriate condition lever to the full aft FUELCUTOFF position, which shuts off the fuel supply.

2-22. POWER LEVERS.

Moving the power levers into reverserange without the engines running mayresult in damage to the reverse linkagemechanism.

Two power levers are located on the controlpedestal (fig. 2-7). These levers regulate power in thereverse, idle, and forward range, and operate so thatforward movement increases engine power. Powercontrol is accomplished through adjustment of theN1 speed governor in the fuel control unit. Power isincreased when N1 RPM is increased. The power

2-23

TM 55-1510-221-10

levers also control propeller reverse pitch. Distinctmovement (pulling up and then aft on the powerlever) by the pilot is required for reverse thrust.Placarding beside the lever travel slots readsPOWER. Upper lever travel range is designatedINCR (increase), supplemented by an arrow point-ing forward. Lower travel range is marked IDLE,LIFT and REVERSE. A placard below the leverslots reads: CAUTION REVERSE ONLY WITHENGINES RUNNING.

2-23. CONDITION LEVERS.

Two condition levers are located on the controlpedestal (fig. 2-7). Each lever starts and stops thefuel supply, and controls the idle speed for itsengine. The levers have three placarded positions:FUEL CUTOFF, LO IDLE, and HIGH IDLE. Inthe FUEL CUTOFF position, the condition levercontrols the cutoff function of its engine-mountedfuel control unit. From LO IDLE to HIGH IDLE,they control the governors of the fuel control unitsto establish minimum fuel flow levels. LO IDLEposition sets the fuel flow rate to attain 52 to 55%(at sea level) minimum N1 and HIGH IDLE posi-tion sets the rate to attain 70% minimum N1. Thepower lever for the corresponding engine can selectN 1 from the respective idle setting to maximumpower. An increase in low idle N1 will be experi-enced at high field elevation.

2-24. FRICTION LOCK KNOBS.

Four friction lock knobs (fig. 2-7) are located onthe control pedestal to adjust friction drag. Oneknob is below the propeller levers, one below thecondition levers, and two under the power levers.When a knob is rotated clockwise, friction restraintis increased opposing movement of the affectedlever as set by the pilot. Counterclockwise rotationof a knob will decrease friction drag thus permittingfree and easy lever movement. Two FRICTIONLOCK placards are located on the pedestal adjacentto the knobs.

2-25. ENGINE FIRE DETECTION SYSTEM.

a. Description. A flame surveillance system isinstalled on each engine to detect external enginefire and provide alarm to the pilot. Both nacelles aremonitored, each having a control amplifier andthree detectors. Electrical wiring connects all sensorsand control amplifiers to DC power and to the cock-pit visual alarm units. In each nacelle, one detectormonitors the forward nacelle, a second monitors theupper accessory area, and a third the lower accessoryarea. Fire emits an infrared radiation that will be

2-24

sensed by the detector which monitors the area oforigin. Radiation exposure activates the relay circuitof a control amplifier which causes signal power tobe sent to cockpit warning systems. An activatedsurveillance system will return to the standby stateafter the fire is out. The system includes a functionaltest switch and has circuit protection through theFIRE 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 warninglight in the appropriate fire control T-handle plac-arded No. 1 FIRE PULL or No. 2 FIRE PULL (fig.2-29). Fire detector circuits are protected by a single5-ampere circuit breaker, placarded FIRE DETR,located on the overhead circuit breaker panel (fig.2-26).

b. Fire Detection System Test Switch. O n erotary switch placarded FIRE PROTECTION TESTon the copilot’s subpanel is provided to test theengine fire detection system. Before checkout, bat-tery power must be on and the FIRE DETR circuitbreaker must be in. Switch position DETR 1, checksthe area forward of the air intake of each nacelle,including circuits to the cockpit alarm and indica-tion devices. Switch position DETR 2, checks thecircuits for the upper accessory compartment ofeach nacelle. Switch position DETR 3, checks thecircuits for the lower accessory compartment of eachnacelle. Each numbered switch position will initiatethe cockpit indications previously described.

c. Erroneous Fire Detection System Indica-tions. During ground test of the engine fire detectionsystem, an erroneous indication of system fault maybe encountered if an engine cowling is not closedproperly, or if the aircraft is headed toward a strongexternal light source. In this circumstance, changethe aircraft heading to enable a valid system check.

2-26. ENGINE FIRE EXTINGUISHER SYSTEM.

a. Description. The fire extinguisher systemutilizes an explosive squib, connected to a valvewhich, when opened, allows the distribution of thepressurized extinguishing agent through a plumbingnetwork of spray nozzles strategically located in thefire zones of the engines.

b . F ire Pul l Handles . The t i re controlT-handles, which are used to arm the extinguishersystem are centrally located on the pilot’s instru-ment panel (fig. 2-29), immediately below theglareshield. These controls receive power from thehot battery bus. The fire detection system will indi-cate an engine fire by illuminating the master faultwarning light on the pilot’s and copilot’s glareshieldand the respective No. 1 FIRE PULL or No. 2 FIRE

PULL lights in the fire control T-handles. Pullingthe fire control T-handle will electrically arm theextinguisher system and close the fuel firewall shut-off valve for that particular engine. This will causethe red light in the PUSH TO EXTINGUISH switchand the respective red No. 1 FUEL PRESS or No. 2FUEL PRESS light on the warning annunciatorpanel to illuminate. Pressing the lens of the PUSHTO EXTINGUISH switch (after lifting one side ofits spring-loaded clear plastic guard) will fire thesquib, expelling all the agent in the cylinder at onetime. The respective yellow caution light, No. 1EXTGH DISCH or No. 2 EXTGH DISCH on thecaution/advisory annunciator panel and the masterfault caution lights on the glareshield will illuminateand remain illuminated, regardless of the masterswitch position, until the squib is replaced.

c. Fire Extinguisher System Test Switch. Arotary test switch, placarded FIRE PROTECTIONTEST, is located on the copilot’s subpanel. The testfunctions, placarded EXTGH - No. 1 - No. 2, arearranged on the left side of the switch and providea test of the pyrotechnic cartridge circuitry. Duringpreflight, the pilot should rotate the test switchthrough the two positions and verify the illumina-tion of the green SQUIB OK light on the PUSH TOEXTINGUISH switch and the corresponding yellowNo. 1 or No. 2 EXTGH DISCH light on the caution/advisory annunciator panel.

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

2-27. OIL SUPPLY SYSTEM.

Maximum allowable oil consumption isone quart per 10 hours of engine opera-tion.

a. The engine oil tank is integral with the air-inlet casting located forward of the accessory gear-box. Oil for propeller operation, lubrication of thereduction gearbox and engine bearings is supplied

TM 55-1510-221-10

by an external line from the high pressure pump.Two scavenge lines return oil to the tank from thepropeller reduction gearbox. A non-congealing exter-nal oil cooler keeps the engine oil temperaturewithin the operating limits. The capacity of eachengine oil tank is 2.3 U.S. gallons. The total systemcapacity for each engine, which includes the oil tank,oil cooler, lines, etc., is 3.5 U.S. gallons. The oillevel is indicated by a dipstick attached to the oilfiller cap. Oil grade, specification and servicingpoints, are described in Section IX, Servicing thewarning annun.

b. The oil system of each engine is coupled toa heat exchanger unit (radiator) of tin-and-tubedesign. These exchanger units are the only airframemounted part of the oil system and are attached tothe nacelles below the engine air intake. Each heatexchanger incorporates a thermal bypass whichassists in maintaining oil at the proper temperaturerange for engine operation.

2-28. ENGINE CHIP DETECTION SYSTEM.

A magnetic chip detector is installed in the bot-tom of each engine nose gearbox to warn the pilot ofoil contamination and possible engine failure. Thesensor is an electrically insulated gap immersed inthe oil, functioning as a normally-open switch. If alarge metal chip or a mass of small particles bridgethe detector gap, a circuit is completed, sending asignal to illuminate a red annunciator panel indica-tor light placarded No. 1 CHIP DETR or No. 2 CHIPDETR and the MASTER WARNING lights. Chipdetector circuits are protected by two 5-ampere cir-cuit breakers, placarded No. 1 CHIP DETR andNo. 2 CHIP DETR on the overhead circuit breakerpanel (fig. 2-26).

2-29. ENGINE IGNITION SYSTEM.

a. Description. The basic ignition system con-sists of a solid state ignition exciter unit, two igniterplugs, two shielded ignition cables, pilot controlledIGNITION AND ENGINE START switches andthe ENG AUTO IGN switch. Placing an IGNITIONAND ENGINE START switch to ON (forward) willcause the respective igniter plugs to spark, ignitingthe fuel/air mixture sprayed into the combustion

Table 2-1. Engine Fire Extinguisher Gage Pressure

TEMP °C -40 -29 -18 -06 04 16 27 38 48

190 220 250 290 340 390 455 525 605PSI to to to to to to to to to

240 275 315 365 420 480 550 635 730

2-25

TM 55-1510-221-10

chamber by the fuel nozzles. The ignition system isactivated for ground and air starts, but is switchedoff after combustion light up.

b. Ignition and Engine Start Switches. Twothree-position toggle switches, placarded IGNITIONAND ENGINE START, are located on the overheadcontrol panel (fig. 2-12). These switches will initiatestarter motoring and ignition in the ON position, orwill motor the engine in the STARTER ONLY posi-tion. The ON switch position completes the startercircuit for engine rotation, energizes the igniter plugsfor fuel combustion, and activates the No. 1 IGNON or No. 2 IGN ON light on the annunciatorpanel. In the center position the switch is OFF. Two5-ampere circuit breakers on the overhead circuitbreaker panel, placarded IGNITOR CONTR No. 1and No. 2, protect ignition circuits. Two 5-amperecircuit breakers on the overhead circuit breakerpanel, placarded START CONTR No. 1 and No. 2,protect starter control circuits (fig. 2-26).

2-30. AUTOIGNITION SYSTEM.

If armed, the autoignition system automaticallyprovides combustion re-ignition of either engineshould accidental flameout occur. The system is notessential to normal engine operation, but is used toreduce the possibility of power loss due to icing or

other conditions. Each engine has a separateautoignition control switch and a green indicatorlight placarded No. 1 IGN ON or No. 2 IGN ON, onthe annunciator panel. Autoignition is accomplishedby energizing the two igniter plugs in each engine.

NOTE

The system should be turned OFF duringextended ground operation to prolong thelife of the igniter plugs.

a. Autoignit ion Switches. Two switches ,located on the overhead control panel (fig. 2-12)placarded ENG AUTO IGN-ARM control theautoignition systems. The ARM position initiates areadiness mode for the autoignition system of thecorresponding engine. The OFF position disarms thesystem. Each switch is protected by a correspondingSTART CONTR No. 1 or No. 2 5-ampere circuitbreaker on the overhead circuit breaker panel (fig.2-26).

b. Autoignition Lights. If an armed autoigni-tion system changes from a ready condition to anoperating condition (energizing the igniter plugs inthe engine) a corresponding green annunciator panellight will illuminate. The annunciator panel light isplacarded No. 1 IGN ON or No. 2 IGN ON and indi-

Figure 2-12. Overhead Control Panel

2-26

cates that the igniters are energized. The autoigni-tion system is triggered from a ready condition to anoperating condition when engine torque drops belowapproximately 20%. Therefore, when an autoignitionsystem is armed, the igniters will be energized con-tinuously during the time when an engine is operat-ing at a level below approximately 20% torque. Theautoignition lights are protected by 5-ampere IGNI-TOR CONTR No. 1 or No. 2 circuit breakers, locatedon the overhead circuit breaker panel (fig. 2-26).

b. Engine Torquemeters. Two torquemeterson the instrument panel indicate torque applied tothe propeller shafts of the respective engines (fig.2-29). Each gage shows torque in percent of maxi-mum using 2 percent graduations and is actuated byan electrical signal from a pressure sensing systemlocated in the respective propeller reduction gearcase. Torquemeters are protected by individual 0.5-ampere circuit breakers placarded TORQUEMETER No. 1 or No. 2 on the overhead circuitbreaker panel (fig. 2-26).

2-31. ENGINE STARTER-GENERATORS.

One starter-generator is mounted on each engineaccessory drive section. Each is able to functioneither as a starter or as a generator. In the starterfunction, 28 volts DC is required to power rotation.In the generator function, each unit is capable of400 amperes DC output. When the starting functionis selected, the starter control circuit receives powerthrough the respective 5-ampere START CONTRcircuit breaker on the overhead circuit breaker panelfrom either the aircraft battery or an external powersource. When the generating function is selected, thestarter-generator provides electrical power. For addi-tional description of the starter-generator system,refer to Section IX.

c. Turbine Tachometers. Two tachometers onthe instrument panel register compressor turbineRPM (N1) for the respective engine (fig. 2-29).These indicators register turbine RPM as a percent-age of maximum gas generator RPM. Each instru-ment is slaved to a tachometer generator attached tothe respective engine.

2-32. ENGINE INSTRUMENTS.

The engine instruments are vertically mountednear the center of the instrument panel (fig. 2-29).

d. Oil Pressure/Oil Temperature Indicators.Two gages on the instrument panel register oil pres-sure in PSI and oil temperature in °C (fig. 2-29). Oilpressure is taken from the delivery side of the mainoil pressure pump. Oil temperature is transmitted bya thermal sensor unit which senses the temperatureof the oil as it leaves the delivery side of the oil pres-sure pump. Each gage is connected to pressure trans-mitters installed on the respective engine. Bothinstruments are protected by 5-ampere circuit break-ers, placarded OIL PRESS and OIL TEMP No. 1 orNo. 2, on the overhead circuit breaker panel (fig.2-26).

a. Turbine Gas Temperature Indicators. TwoTGT gages on the instrument panel are calibrated indegrees Celsius (fig. 2-29). Each gage is connected tothermocouple probes located in the hot gasesbetween the turbine wheels. The gages register thetemperature present between the compressor turbineand power turbine for the corresponding engine.

e. Fuel Flow Indicators. Two gages on theinstrument panel (fig. 2-29) register the rate of flowfor consumed fuel as measured by sensing units cou-pled into the fuel supply lines of the respectiveengines. The fuel flow indicators are calibrated inincrements of hundreds of pounds per hour. Bothcircuits are protected by 0.5-ampere circuit breakersplacarded FUEL FLOW No. 1 or No. 2, on the over-head circuit breaker panel (fig. 2-26).

Section IV. FUEL SYSTEM

2-33. FUEL SUPPLY SYSTEM.

The engine fuel supply system (fig. 2-13) consistsof two identical systems sharing a common fuelmanagement panel (fig. 2-14) and fuel crossfeedplumbing (fig. 2-15). Each fuel system consists offive interconnected wing tanks, a nacelle tank, andan auxiliary inboard fuel tank. A fuel transfer pumpis located within each auxiliary tank. Additionally,

the system has an engine-driven boost pump, astandby fuel pump located within each nacelle tank,a fuel heater (engine oil-to-fuel heat exchanger unit),a tank vent system, a tank vent heating system andinterconnecting wiring and plumbing. Refer to Sec-tion IX for fuel grades and specifications. Fuel tankcapacity is shown in table 2-2. Gravity feed fuel flowis shown in figure 2-16.

TM 55-1510-221-10

2-27

TM 55-1510-221-10

Figure 2-13. Fuel System Schematic

2-28

TM 55-1510-221-10

1. STANDBY PUMP switch (#1 engine)2. FUEL quantity indicator (#1 engine)3. FUEL QUANTITY gaging system control switch4. FUEL quantity indicator (#2 engine)5. STANDBY PUMP switch (#2 engine)6. AUX TRANSFER OVERRIDE switch (#2 engine)7. CROSSFEED valve switch8. AUX TRANSFER OVERRIDE switch (#1 engine)

Figure 2-14. Fuel Management Panel

a. Engine Driven Boost Pumps.

Engine operation using only the engine-driven primary (high pressure) fuel pumpwithout standby pump or engine-drivenboost pump fuel pressure is limited to 10cumulative hours. This condition is indi-cated by illumination of either the No. 1or No. 2 FUEL PRESS warning annuncia-tor lights and the simultaneous illumina-tion of both MASTER WARNING lights.Refer to Chapter 9. All time in this cate-gory shall be entered on DA Form2408-13 for the attention of maintenancepersonnel.

A gear-driven boost pump, mounted on eachengine supplies fuel under pressure to the inlet ofthe engine-driven primary high-pressure pump forengine starting and all normal operations. Either theengine-driven boost pump or standby pump is capa-ble of supplying sufficient pressure to the engine-driven primary high-pressure pump and thus main-tain normal engine operation.

AP 006447

b. Standby Fuel Pumps. A submerged, electri-cally-operated standby fuel pump, located withineach nacelle tank, serves as a backup unit for theengine-driven boost pump. The standby pumps areswitched off during normal system operations. Astandby fuel pump will be operated during crossfeedoperation to pump fuel from one system to theopposite engine. The correct pump is automaticallyselected when the CROSSFEED switch is activated.Each standby fuel pump has an inertia switchincluded in the power supply circuit. When sub-jected to a 5 to 6 G shock loading, as in a crash situ-ation, the inertia switch will remove electrical powerfrom the standby fuel pumps. The standby fuelpumps are protected by two 10-ampere circuitbreakers placarded STANDBY PUMP No. 1 orNo. 2, located the overhead circuit breaker panel (fig.2-26) and four 5-ampere circuit breakers (2 each inparallel) on the hot battery bus.

c. Fuel Transfer Pumps. The auxiliary tankfuel transfer system automatically transfers the fuelfrom the auxiliary tank to the nacelle tank withoutpilot action. Motive flow to a jet pump mounted inthe auxiliary tank sump is obtained from the enginefuel plumbing system downstream from the enginedriven boost pump and routed through the transfercontrol motive flow valve. The motive flow valve isenergized to the open position by the control system

2-29

TM 55-1510-221-10

Figure 2-15. Crossfeed Fuel Flow

2-30

TM 55-1510-221-10

Figure 2-16. Gravity Feed Fuel Flow

2-31

TM 55-1510-221-10

Table 2-2. Fuel Quantity Data

T A N K S NUMBER GALLONS **POUNDS

Wing Tanks 5 135 877.5LEFT ENGINE Nacelle Tank 1 57 370.5

Auxiliary Tank 1 79 513.5

Wing Tanks 5 135 884.0RIGHT ENGINE Nacelle Tank 1 57 370.5

Auxiliary Tank 1 79 513.5

*TOTALS 14 542 3523.0

* Unusable fuel quantity and weight (4 gallons, 26 pounds) not included in totals.* * Fuel weight is based on standard day conditions at 6.5 pounds per U.S. gallon. Total fuel system capacity is 542

gallons (usable).

to transfer auxiliary fuel to the nacelle tank to beconsumed by the engine during the initial portion of theflight. When an engine is started, pressure at the enginedriven boost pump closes a pressure switch which, aftera 30 to 50 second tie delay to avoid depletion of fuelpressure during starting, energizes the motive flowvalve. When auxiliary fuel is depleted, a low level floatswitch de-energizes the motive flow valve after a 30 to 60second time delay provided to prevent cycling of themotive flow valve due to sloshing fuel. In the event of afailure of the motive flow valve or the associated controlcircuitry, the loss of motive flow pressure when there isstill fuel remaining in the auxiliary fuel tank is sensed bya pressure switch and float switch, respectively, whichilluminates a caution annunciator light placarded No. 1NO FUEL XFR or No. 2 NO FUEL XFR. 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 tanksare empty. A manual override is incorporated as abackup for the automatic transfer system. This isinitiated by placing the AUX TRANSFER switch, locatedon the FUEL management panel to the OVERRIDEposition. This will energize the transfer control motiveflow valve. The transfer systems are protected by5-ampere circuit breakers placarded AUXILIARYTRANSFER No. 1 or No. 2, located on the overheadcircuit breaker panel (fig. 2-26), 2.0 inches high.

NOTE

In turbulence or during maneuvers, the NOFUEL XFR lights may momentarilyilluminate after the auxiliary fuel hascompleted transfer.

d. Fuel Gaging System. The total fuel quantity inthe left or right main system or left or right auxiliarytank is measured by a capacitance type fuel gaging

2-32 Charge 4

system. Two fuel gages, one for the left and one for theright fuel system, read fuel quantity in pounds. Refer toSection IX for fuel capacities and weights. A maximumof 3% error may be encountered in each system.However, the system is compensated for fuel densitychanges due to temperature excursions. In addition tothe fuel gages, yellow No. 1 NAC LOW or No. 2 NACLOW lights on the caution/advisory annunciatorpanuminate when there is approximately 20 minutes offuel per engine remaining (on standard day, at sea level,normal cruise power consumption rate). The fuel gagingsystem is protected by individual 5-ampere circuitbreakers placarded QTY IND and QTY WARN No. 1 orNo. 2, located on the overhead circuit breaker panel (fig.2-26). A mechanical spiral float gage is installed in eachauxiliary fuel tank to provide an indication of fuel levelwhen servicing the tank. The gage is installed on theauxiliary fuel tank cover, adjacent to the filler neck (fig.2-13). A small sight window in the upper wing skinpermits observation of the gage.

e. Fuel Management Panel. The fuel managementpanel is located on the cockpit overhead between thepilot and copilot. It contains the fuel gages, standby fuelpump switches, the crossfeed valve switch and a fuelgaging system control switch and transfer controlswitches are also installed.

(1.) Fuel gaging system control switch. A switch onthe fuel management panel (fig. 2-14) placarded FUELQUANTITY, MAIN - AUXILIARY, controls the fuelgaging system. When in the MAIN position the fuelgages 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. T w oswitches, placarded STANDBY PUMP - ON locatedon the fuel management panel (fig. 2-14) control asubmerged fuel pump located in the correspondingnacelle tank. During normal aircraft operation bothswitches are off so long as the engine-driven boostpumps function and during crossfeed operation. Theloss of fuel pressure, due to failure of an enginedriven boost pump will illuminate the MASTERWARNING lights on the glareshield and will illumi-nate the No. 1 FUEL PRESS or No. 2 FUEL PRESSon the warning annunciator panel. Turning ON theSTANDBY PUMP will extinguish the FUEL PRESSlights. The MASTER WARNING lights must bemanually cleared.

NOTE

Both standby pump switches shall be offduring crossfeed operation.

(3.) Fuel transfer control switches. Twoswitches on the fuel management panel (fig. 2-14),placarded AUX TRANSFER OVERRIDE - AUTOcontrol operation of the fuel transfer pumps Duringnormal operation both switches are in AUTO whichallows the system to be automatically actuated byfuel flow to the engine. If either transfer system failsto operate, the fault condition is indicated by twoilluminated MASTER CAUTION lights on theglareshield and a steadily illuminated yellow No. 1NO FUEL XFR or No. 2 NO FUEL XFR light onthe caution annunciator panel.

(4.) Fuel crossfeed switch. The fuel cross-feed valve is controlled by a 3-position switch (fig.2-14), located on the fuel management panel, plac-arded CROSSFEED - OFF. Under normal flightconditions the switch is left in the OFF position.During emergency single engine operation, it maybecome necessary to supply fuel to the operativeengine from the fuel system on the opposite side.The crossfeed system is placarded for fuel selectionwith a simplified diagram on the overhead fuel con-trol panel. Place the standby fuel pump switches inthe off position when crossfeeding. A lever lockswitch, placarded CROSSFEED, is moved from thecenter OFF position to the left or to the right,depending on direction of fuel flow. This opens thecrossfeed valve and energizes the standby pump onthe side from which crossfeed is desired. Duringcrossfeed operation with firewall fuel valve closed,auxiliary tank fuel will not crossfeed. When thecrossfeed mode is energized, a green FUEL CROSS-FEED light on the caution/advisory panel will illu-minate. Crossfeed system operation is described inChapter 9. The crossfeed valve is protected by a5-ampere circuit breaker placarded CROSSFEED

TM 55-1510-221-10

located on the overhead circuit breaker panel (fig.2-26).

f. Firewall Shutoff Valves.

Do not use the fuel firewall shutoff valveto shut down an engine, except in anemergency. The engine-driven high-pressure fuel pump obtains essentiallubrication from fuel flow. When anengine is operating, this pump may beseverely damaged (while cavitating) if thefirewall valve is closed before the condi-tion lever is moved to the FUEL CUT-OFF 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 arepulled out. The firewall shutoff valves receive elec-trical power from the main buses and also from thehot battery bus which is connected directly to thebattery. The valves are protected by circuit breakersplacarded FIREWALL VALVE No. 1 or No. 2 on theoverhead circuit breaker panel (fig. 2-26) andFIREWALL SHUTOFF No. 1 or No. 2 on the hotbattery bus circuit breaker board.

g. Fuel Tank Sump Drains. A sump drainwrench is provided in the aircraft loose tools to sim-plify draining a small amount of fuel from the sumpdrain.

(1.) There are five sump drains and onefilter drain in each wing (Table 2-3).

(2.) An additional drain for the extendedrange fuel system line extends through the bottom ofthe wing center section adjacent to the fuselage.Anytime the extended range system is in use, a partof the preflight inspection would consist of draininga small amount of fuel from this drain to check forfuel contamination. Whenever the extended rangesystem is removed from the aircraft and the fuel lineis capped off in the fuselage, the remaining fuel inthe line shall be drained.

h. Fuel Drain Collector System. Each engine isprovided with a fuel drain collector system to returnfuel dumped from the engine during clearing andshutdown operations back into its respective nacelletank. The system draws power from the No. 4 feederbus. Fuel transfer is completely automatic. Fuelfrom the engine flow divider drains into a collectortank mounted below the aft engine accessory sec-

2-33

TM 55-1510-221-10

Table 2-3. Fuel Sump Drain Locations

NUMBER DRAINS LOCATION

1 Leading Edge Tank Outboard of nacelle, underside of wing1 Integral Tank Underside of wing, forward of aileron1 Firewall Fuel Filter Underside of cowling forward of firewall1 Sump Strainer Bottom center of nacelle forward of wheel well

1 Gravity Feed Line Aft of wheel well1 Auxiliary Tank At wing root, just forward of the flap

tion. An internal float switch actuates an electricscavenger pump which delivers the fuel to the fuelpurge line just aft of the fuel purge shutoff valve. Acheck valve in the line prevents the backflow of fuelduring engine purging. The circuit breaker for bothpumps is located in the fuel section of the overheadcircuit breaker panel; placarded SCAVENGERPUMP. A vent line, plumbed from the top of thecollector tank, is routed through an inline flamearrestor and then downward to a drain manifold onthe underside of the nacelle.

fer of auxiliary fuel, which is automatically con-trolled, the nacelle tanks are maintained full. Acheck valve in the gravity feed line from the out-board wing prevents reverse fuel flow. Normal grav-ity transfer of the main wing fuel into the nacelletanks will begin when auxiliary fuel is exhausted.The system will gravity feed fuel only to its respec-tive nacelle tank, i.e. left or right (fig. 2-16). Fuelwill not gravity feed through the crossfeed system.

i. Fuel Vent System. Each fuel system isvented through two ram vents located on the under-side of the wing adjacent to the nacelle. To preventicing of the vent system, one vent is recessed intothe wing and the backup vent protrudes out fromthe wing and contains a heating element. The ventline at the nacelle contains an inline flame arrestor.

j. Engine Oil-to-Fuel Heat Exchanger. A nengine oil-to-fuel heat exchanger, located on eachengine accessory case, operates continuously andautomatically to heat the fuel delivered to the enginesufficiently to prevent the freezing of any waterwhich it might contain. The temperature of thedelivered fuel is thermostatically regulated to remainbetween 21°C and 32°C.

b. Operation With Failed Engine-Driven BoostPump or Standby Pump. Two pumps in each fuelsystem provide inlet head pressure to the engine-driven primary high-pressure fuel pump. If crossfeedis used, a third pump, the standby fuel pump fromthe opposite system, will supply the required pres-sure. Operation under this condition will result in anunbalanced fuel load as fuel from one system will besupplied to both engines while all fuel from the sys-tem with the failed engine driven and standby boostpumps will remain unused. A triple failure, which ishighly unlikely, would result in the engine drivenprimary pump operating without inlet head pres-sure. Should this situation occur, the affected enginecan continue to operate from its own fuel supply onits engine-driven primary high-pressure fuel pump.

2-34. FUEL SYSTEM MANAGEMENT. 2-35. FERRY FUEL SYSTEM.

a. Fuel Transfer System. When the auxiliarytanks are filled, they will be used first. During trans-

Provisions are installed for connection to longrange fuel cells.

Section V. FLIGHT CONTROLS

2-36. DESCRIPTION.

The aircraft’s primary flight control systemsconsist of conventional rudder, elevator and aileroncontrol surfaces. These surfaces are manually oper-ated from the cockpit through mechanical linkageusing a control wheel for the ailerons and elevators,and adjustable rudder/brake pedals for the rudder.Both the pilot and copilot have flight controls. Trim

control for the rudder, elevator and ailerons isaccomplished through a manually actuated cable-drum system for each set of control surfaces. Theautopilot has provisions for controlling the positionof the ailerons, elevators, and rudder. Chapter 3describes the operation of the autopilot system.

2-34

2-37. CONTROL WHEELS

Elevator and aileron control surfaces are oper-ated by manually actuating either the pilot’s or copi-lot’s control wheel. Switches are installed in the out-board grip of each wheel to operate the elevator trimtabs. A microphone switch, a chaff dispense switch,and an autopilot/yaw damp/electric trim disconnectswitch are also installed in the outboard grip of eachwheel. A transponder ident switch is installed on topof the inboard grip of each control wheel. These con-trol wheels (fig. 2-17) are installed - on each side ofthe instrument subpanel. A manually wound 8-dayclock is installed in the center of the pilot’s controlwheel, and a digital electric clock is installed in thecenter of the copilot’s control wheel. A map lightswitch, and a pitch synchronization and controlwheel steering switch are mounted adjacent to theclock in each control wheel.

2-38. RUDDER SYSTEM.

a. Rudder Pedals. Aircraft rudder control andnose wheel steering is accomplished by actuation ofthe rudder pedals from either pilot’s or copilot’s sta-tion (fig. 2-8). The rudder pedals may be individu-ally adjusted in either a forward or aft position toprovide adequate leg room for the pilot and copilot.Adjustment is accomplished by depressing the leveralongside the rudder pedal arm and moving thepedal forward or aft until the locking pin engages inthe selected position.

b. Yaw Damp System. A yaw damp system isprovided to aid the pilot in maintaining directionalstability and increase ride comfort. The system maybe used at any altitude and is required for flightabove 17,000 feet. It must be deactivated for takeoffand landing. The yaw damp system is a part of theautopilot. Operating instructions for this system arecontained in Chapter 3. The system is controlled bya YAW DAMP switch adjacent to the ELEV TRIMswitch on the pedestal extension.

c. Rudder Boost System. A rudder boost sys-tem is provided to aid the pilot in maintainingdirectional control resulting from an engine failureor a large variation of power between the engines.Incorporated in the rudder cable system are twopneumatic rudder boosting servos which actuate thecables to provide rudder pressure to help compen-sate for asymmetrical thrust.

(1.) During operation, a differential pres-sure valve accepts bleed air pressure from eachengine. When the pressure varies between the bleedair systems, the shuttle in the differential pressurevalve moves toward the low pressure side. As thepressure difference reaches a preset tolerance, a

TM 55-1510-221-10

switch closes on the low pressure side which acti-vates the rudder boost system. This system isdesigned only to help compensate for asymmetricalthrust. Appropriate trimming is to be accomplishedby the pilot. Moving either or both of the bleed airvalve switches on the overhead control panel toPNEU & ENVIRO - OFF position will disengage therudder boost system.

NOTE

Condition levers must be in LOW IDLEposition to perform rudder boost check.

(2.) The system is controlled by a switchlocated on the extended pedestal placarded RUD-DER BOOST (on) - OFF (fig. 2-7), and is to beturned on before flight. A preflight check of the sys-tem can be performed during the run-up by retard-ing the power on one engine to idle and advancingpower on the opposite engine until the power differ-ence between the engines is great enough to activatethe switch to turn on the rudder boost system.Movement of the appropriate rudder pedal (leftengine idling, right rudder pedal moves forward) willbe noted when the switch closes, indicating the sys-tem is functioning properly for low engine power onthat side. Repeat the check with opposite power set-tings to check for movement of the opposite rudderpedal. The system is protected by a 5-ampere circuitbreaker placarded RUDDER BOOST, located onthe overhead circuit breaker panel (fig. 2-26).

NOTE

With brake deice on, rudder boost may beinoperative.

2-39. FLIGHT CONTROLS LOCK.

Remove control locks before towing theaircraft or starting engines. Serious dam-age could result in the steering linkage iftowed by a tug with the rudder lockinstalled.

Positive locking of the rudder, elevator and aile-ron control surfaces, and engine controls (powerlevers, propeller levers, and condition levers) is pro-vided by a removable lock assembly (fig. 2-18) con-sisting of two pins, and an elongated U-shaped strapinterconnected by a chain. Installation of the controllocks is accomplished by inserting the U-shapedstrap around the aligned control levers from the

2-35

TM 55-1510-221-10

(Pilot)

1. Microphone, intercom, transmit switch2. Trim, autopilot, yaw damp disconnect switch3. Pitch-trim switches4. Chaff dispense switch5. Transponder ident switch6. Map light7. Eight day clock8. Pitch synchronization and control wheel steering switch

(Copilot)

AP 010329

1. Microphone, intercom, transmit switch2. Chaff dispense switch3. Pitch-trim switches4. Trim, autopilot, yaw damp diconnect switch5. Transponder ident switch6. Pitch synchronization and control wheel steering switch7. Digital clock8. Map light

Figure 2- 17. Control Wheels

2-36

TM 55-1510-221-10

APOO5445

Figure 2- 18. Control Locks

copilot’s side; then the aileron/elevator locking pinis inserted through a guide hole in the top of thepilot’s control column assembly, thus locking thecontrol wheel. The rudder is held in a neutral posi-tion by an L-shaped pin which is installed througha guide hole in the floor aft of the pilots rudder ped-als. The rudder pedals must be centered to align thehole in the rudder bellcrank with the guide hole inthe floor. Remove the locks in reverse order, i.e.,rudder pin, control column pin, and power controlclamp.

2-40. TRIM TABS.

Trim tabs are provided for all flight control sur-faces. These tabs are manually activated, and aremechanically controlled by a cable-drum and jack-screw actuator system, except the right aileron tabwhich is of the fixed bendable type. Elevator andaileron trim tabs incorporate neutral, non-servoaction, i.e., as the elevators or ailerons are displacedfrom the neutral position, the trim tab maintains an“as adjusted” position. The rudder trim tab incorpo-rates anti-servo action, i.e., as the rudder is dis-placed from the neutral position the trim tab movesin the same direction as the control surface. Thisaction increases control pressure as rudder isdeflected from the neutral position.

a. Elevator Trim Tab Control. The elevatortrim tab control wheel placarded ELEVATOR TAB-DOWN, UP, is on the left side of the control pedes-tal and controls a trim tab on each elevator (fig.2-7). The amount of elevator tab deflection, indegrees from a neutral setting, is indicated by a posi-tion arrow.

b. Electric Elevator Trim. The electric eleva-tor trim system is controlled by an ELEV TRIM -ON - OFF/RESET switch located on the pedestal,dual element thumb switches on the control wheels,a trim disconnect switch on each control wheel anda circuit breaker on the overhead circuit breakerpanel. The ON - OFF/RESET switch must be in theON position to operate the system. The dual ele-ment thumb switch is moved forward for trimmingnose down, aft for nose up, and when releasedreturns to the center (off) position. Any activation ofthe trim system through the copilot’s trim switchcan be over ridden by activation of the pilot’sswitch. Operating the pilot’s and copilot’s switchesin opposing directions simultaneously results in thepilot having priority.

A preflight check of the switches should beaccomplished before flight by moving the switchesindividually on both control wheels. No one switchalone should operate the system; operation of eleva-

2-37

TM 55-1510-221-10

tor trim should occur only by movement of pairs ofswitches. The trim system disconnect is a bi-level,pushbutton, momentary type switch, located on theoutboard grip of each control wheel. Depressing theswitch to the first of two levels disconnects theautopilot and yaw damp system, and the secondlevel disconnects the electric trim system. The sys-tem can be reset by moving the ELEV TRIM switchtoggle on the pedestal (fig. 2-7) to OFF RESET posi-tion, then back to ELEV TRIM (on) again.

c. Aileron Trim Tab Control. The aileron trimtab control, placarded AILERON TAB - LEFT,RIGHT, is on the control pedestal and will adjustthe left aileron trim tab only (fig. 2-7). The amountof aileron tab deflection, from a neutral setting, asindicted by a position arrow, is relative only and isnot in degrees. Full travel of the tab control movesthe trim tab 7-1/2 degrees up and down.

d. Rudder Trim Tab Control. The rudder trimtab control knob, placarded RUDDER TAB -LEFT,RIGHT, is on the control pedestal, and controlsadjustment of the rudder trim tab (fig. 2-7). Theamount of rudder tab deflection, in degrees from aneutral setting, is indicated by a position arrow.

2-41. 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.During extension, or retraction, the flaps are oper-ated as a single unit, each section being actuated bya separate jackscrew actuator. The actuators aredriven through flexible shafts by a single, reversibleelectric motor. Wing flap movement is indicated inpercent of travel by a flap position indicator on theforward control pedestal. Full flap extension andretraction time isflap control switch

approximately 11 seconds. Theis located on the control pedestal.

b. Wing Flap Position Indicator. Flap positionin percent of travel from “0“ percent (UP) to 100percent (DOWN), is shown on an indicator, plac-arded FLAPS located on the control pedestal (fig.2-7). The approach and full down or extended flapposition is 14 and 34 degrees, respectively. The flapposition indicator is protected by a 5-ampere circuitbreaker, placarded FLAP CONTR, located on theoverhead circuit breaker panel (fig. 2-26).

Section VI. PROPELLERS

2-42. DESCRIPTION.

A three-blade aluminum propeller is installed oneach engine. The propeller is of the full feathering,constant speed, counterweighted, reversible type,controlled by engine oil pressure through singleaction, engine driven propeller governors. The pro-peller is flange-mounted to the engine shaft. Centrif-ugal counterweights, assisted by a feathering spring,move the blades toward the low RPM (high pitch)position and into the feathered position. Governorboosted engine oil pressure moves the propeller to

No emergency wing flap actuation system is pro-vided. With flaps extended beyond the APPROACHposition, the landing gear warning horn will sound,unless the landing gear is down and locked. The cir-cuit is protected by a 20-ampere circuit breaker,placarded FLAP MOTOR, located on the overheadcircuit breaker panel (fig. 2-26).

a. Wing Flap Control Switch. Flap operationis controlled by a three-position switch with a flap-shaped handle on the control pedestal (fig. 2-7). Thehandle of this switch is placarded FLAP and switchpositions are placarded: FLAP - UP, APPROACH,and DOWN. The amount of downward extension ofthe flaps is established by position of the flap switch,and is as follows: UP - 0%, APPROACH - 40%, andDOWN -100%. Limit switches, mounted on theright inboard flap, control flap travel. The flap con-trol switch, limit switch, and relay circuits are pro-tected by a 5-ampere circuit breaker, placardedFLAP CONTR located on the overhead circuitbreaker panel (fig. 2-26). Flap positions between UPand APPROACH cannot be selected. For intermedi-ate flap positions between APPROACH andDOWN, the APPROACH position acts as an offposition. To return the flaps to any position betweenfull DOWN and APPROACH, place the flap switchto 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 to5 degrees out of phase with the other, a safety mech-anism is provided to discontinue power to the flapmotor.

the high RPM (low pitch) hydraulic stop and reverseposition. The propellers have no low RPM (highpitch) stops; this allows the blades to feather afterengine shutdown. Low pitch propeller position isdetermined by the low pitch stop which is a mechan-ically actuated, hydraulic stop. Beta and reverseblade angles are controlled by the power levers inthe beta and reverse range.

2-38

TM 55-1510-221-10

2-43. FEATHERING PROVISIONS.

Both manual and automatic propeller featheringsystems are provided. Manual feathering is accom-plished by pulling the corresponding propeller leveraft past a friction detent. To unfeather, the propellerlever is pushed forward into the governing range. Anautomatic feathering system, will sense loss oftorque and will feather an unpowered propeller.Feathering springs will feather the propeller when itis not turning.

a. Automatic Feathering. The automaticfeathering system provides a means of immediatelydumping oil from the propeller servo to enable thefeathering spring and counterweights to start feath-ering action of the blades in the event of an enginefailure. Although the system is armed by a switch onthe overhead control panel, placarded AUTOFEA-THER - ARM - OFF - TEST, the completion of thearming phase occurs when both power levers areadvanced above 90% N1 - at which time both indica-tor lights on the caution/advisory annunciator panelindicate a fully armed system. The annunciatorpanel lights are green and are placarded No.1AUTOFEATHER (left engine) and No.2AUTOFEATHER (right engine). The system willremain inoperative as long as either power lever isretarded below 90% N1 position, unless TEST posi-tion of the AUTOFEATHER SWITCH is selected todisable the power lever limit switches. The system isdesigned for use only during takeoff and landing andshould be turned off when establishing cruise climb.During takeoff or landing, should the torque foreither engine drop to an indication between 16 -21%, the autofeather system for the opposite enginewill be disarmed. Disarming is confirmed when theNo.1 AUTOFEATHER or No.2 AUTOFEATHERannunciator light of the opposite engine becomesextinguished. If torque drops further, to a readingbetween 9 and 14%, oil is dumped from the servo ofthe affected propeller allowing a feathering spring tomove the blades into the feathered position. Feath-ering also causes the No.1 AUTOFEATHER orNo.2 AUTOFEATHER annunciator light of thefeathered propeller to extinguish. At this time, boththe No.1 AUTOFEATHER and No.2 AUTOFEA-THER lights are extinguished, the propeller of thedefective engine has feathered, and the propeller ofthe operative engine has been disarmed from theautofeathering capability. Only manual featheringcontrol remains for the second propeller.

b. Propeller Autofeather Switch. Autofeather-ing is controlled by an AUTOFEATHER switch onthe overhead control panel (fig. 2-12). The three-position switch is placarded ARM, OFF and TEST,and is spring loaded from TEST to OFF. The ARMposition is used only during takeoff and landing.The TEST position of the switch, enables the pilot

to check readiness of the autofeather systems, below88% to 92% Nl , and is for ground checkout purposesonly.

c. Autofeather Lights. Two green lights on thecaution/advisory annunciator panel are placardedAUTOFEATHER No. 1 and AUTOFEATHERNo.2. When illuminated, the lights indicate that theautofeather system is armed. Both lights will beextinguished if either propeller has been autofea-thered or if the system is disarmed by retarding apower lever. Autofeather circuits are protected byone 5-ampere circuit breaker placarded AUTOFEATHER, located on the overhead circuit breakerpanel (fig. 2-26).

2-44. PROPELLER GOVERNORS.

Two governors, a constant speed governor, andan overspeed governor, control propeller RPM. Theconstant speed governor, mounted on top of thereduction housing, controls the propeller through itsentire range. The propeller control lever operates thepropeller by means of this governor. If the constantspeed governor should malfunction and requestmore than 2000 RPM, the overspeed governor cutsin at 2080 RPM and dumps oil from the propellerto keep the RPM from exceeding approximately2080 RPM. A solenoid, actuated by the PROP GOVTEST switch located on the overhead control panel(fig. 2-12), is provided for resetting the overspeedgovernor to approximately 1830 to 1910 RPM fortest purposes. If the propeller sticks or moves tooslowly during a transient condition causing the pro-peller governor to act too slowly to prevent an over-speed condition, the power turbine governor, con-tained within the constant speed governor housing,acts as a fuel topping governor. When the propellerreaches 106% of N2 RPM, the fuel topping governorlimits the fuel flow to the gas generator, reducing N1RPM, which in turn prevents the propeller RPMfrom exceeding approximately 2120 RPM. Duringoperation in the reverse range, the power turbinegovernor is reset to approximately 95% of propellerRPM before the propeller reaches a negative pitchangle. This insures that the engine power is limitedto maintain a propeller RPM of somewhat less thanthat of the constant speed governor setting. Theconstant speed governor therefore, will always sensean underspeed condition and direct oil pressure tothe propeller servo piston to permit propeller opera-tion in beta and reverse ranges.

2-45. PROPELLER TEST SWITCHES.

Two two-position switches on the overhead con-trol panel (fig. 2-12), are provided for operationaltesting of the propeller systems. Placarding above

2-39

TM 55-1510-221-10

the switches reads PROP GOV TEST. Each switchcontrols test circuits for the corresponding propeller.In the test position, the switches are used to test thefunction of the corresponding overspeed governor.Refer to Chapter 8, for test procedure. Propeller testcircuits are protected by one 5-ampere circuitbreaker placarded PROP GOV located on the over-head circuit breaker panel (fig. 2-26).

2-46. PROPELLER SYNCHROPHASER.

a. Operation. The propeller synchrophaserautomatically matches the RPM of the right propel-ler (slave propeller) to that of the left propeller (mas-ter propeller) and maintains the blades of one pro-peller at a predetermined relative position with theblades of the other propeller. To prevent the rightpropeller from losing excessive RPM if the left pro-peller is feathered while the synchrophaser is on, thesynchrophaser has a limited range of control fromthe manual governor setting. Normal governor oper-ation is unchanged but the synchrophaser will con-tinuously monitor propeller RPM and reset the gov-ernor as required. A magnetic pickup mounted ineach propeller overspeed governor and adjacent toeach propeller deice brush block transmits electricpulses to a transistorized control box installed for-ward of the pedestal. The right propeller RPM andphase will automatically be adjusted to correspondto the left. To change RPM, adjust both propellercontrols at the same time. This will keep the rightgovernor setting within the limiting range of the leftpropeller. If the synchrophaser is on but is unable toadjust to the right propeller to match the left, theactuator has reached the end of its travel. Torecenter, turn the switch off, synchronize the propel-lers manually, and turn the switch back on.

b. Control Box. The control box converts anypulse rate differences into correction commands,which are transmitted to a stepping type actuatormotor mounted on the right engine cowl forwardsupport ring. The motor then trims the right propel-ler governor through a flexible shaft and trimmerassembly to exactly match the left propeller. Thetrimmer, installed between the governor control armand the control cable, screws in or out to adjust thegovernor while leaving the control lever setting cons-tant. A toggle switch installed adjacent to the syn-chrophaser turns the system on. With the switch off,the actuator automatically runs to the center of itsrange of travel before stopping to assure normalfunction when used again. To operate the system,synchronize the propeller in the normal manner andturn the synchrophaser on. The system is designedfor in-flight operations and is placarded to be off fortake-off and landing. Therefore, with the system on

and the landing gear extended, the master cautionlights will illuminate and a yellow light on the cau-tion/advisory annunciator panel, PROP SYNC ON,will illuminate.

c. Synchroscope. The propeller synchroscope,provides an indication of synchronization of thepropellers. If the right propeller is operating at ahigher RPM than the left, the face of the synchro-scope, a black and white cross pattern, spins in aclockwise rotation. Left, or counterclockwise, rota-tion indicates a higher RPM of the left propeller.This instrument aids the pilot in obtaining completesynchronization of propellers. The system is pro-tected by a 5-ampere circuit breaker placardedPROP SYNC, located on the overhead circuitbreaker panel (fig. 2-26).

2-47. PROPELLER LEVERS.

Two propeller levers on the control pedestal (fig.2-7), placarded PROP, are used to regulate propellerspeeds. Each lever controls a primary governor,which acts to regulate propeller speeds within thenormal operation range. The full forward position ofthe levers is placarded TAKEOFF, LANDING ANDREVERSE - and also HIGH RPM. The full aft posi-tion of the levers is placarded FEATHER. When alever is placed at HIGH RPM, the propeller mayattain a static RPM of 2,000 depending upon powerlever position. As a lever is moved aft, passingthrough the propeller governing range, but stoppingat the feathering detent, propeller RPM will corre-spondingly decrease to the lowest limit. Moving apropeller lever aft past the detent into FEATHERwill feather the propeller.

2-48. PROPELLER REVERSING.

Do not move the power levers intoreverse range without the engine running.Damage to the reverse linkage mecha-nisms will occur.

Propeller reversing on unimproved sur-faces should be accomplished carefully toprevent propeller erosion from reversedairflow and, in dusty conditions, to pre-vent obscuring the operator’s vision.

2-40

TM 55-1510-221-10

To prevent an asymmetrical thrust condi-tion, propeller levers must be in HIGHRPM position prior to propeller revers-ing.

The propeller blade angle may be reversed toshorten landing roll. To reverse, propeller leversmust be positioned at HIGH RPM (full forward),and the power levers are lifted up to pass over theIDLE detent, then pulled aft into REVERSE setting.One yellow caution light, placarded REV NOTREADY, on the caution/advisory annunciator panel,

alerts the pilot not to reverse the propellers. Thislight illuminates only when the landing gear handleis down, and if propeller levers are not at HIGHRPM (full forward). This circuit is protected by a5-ampere circuit breaker, placarded LANDINGGEAR RELAY, located on the overhead circuitbreaker panel (fig. 2-26).

2-49. PROPELLER TACHOMETERS.

Two tachometers on the instrument panel regis-ter propeller speed in hundreds of RPM (fig. 2-29).Each indicator is slaved to a tachometer generatorunit attached to the corresponding engine.

Section VII. UTILITY SYSTEMS

2-50. DEFROSTING SYSTEM.

a. Description. The defrosting system is anintegral part of the heating and ventilation system.The system consists of two warm air outlets con-nected by ducts to the heating system. One outlet isjust below the pilot’s windshield and the other isbelow the copilot’s windshield. A push-pull control,placarded DEFROST AIR, on the pilot’s subpanel,manually controls airflow to the windshield. Whenpulled out, defrosting air is ducted to the wind-shield. As the control is pushed in, there is a corre-sponding decrease in airflow.

b. Automatic Operation.

1. Vent blower switches - As required.

2 . Cabin temperature mode se lectorswitch - AUTO.

3. Cabin temperature control rheostat -As required.

4. Cabin air, copilot air, pilot air, anddefrost air controls - As required.

c. Manual Operation.

1. Pilot air, copilot air - IN.

2. Cabin air and defrost air controls - Out

3. Cabin temperature mode se lectorswitch - MAN HEAT.

4. Cold air outlets - As required.

5 . Manual temperature swi tch - Asrequired.

d. Manual Operation. If the automatic tem-perature control should fail to operate, the tempera-ture (of defrost air and cabin air) may be controlledmanually by setting the CABIN TEMP MODE con-trol switch to MANUAL COOL position, then usingthe MANUAL TEMP DECREASE-INCREASEswitch to set the desired temperature. This control islocated on the overhead control panel (fig. 2-l 2).

2-51. SURFACE DEICING SYSTEM.

a. Description. Ice accumulation is removedfrom each inboard and outboard wing leading edge,both horizontal stabilizers, the taillets, and certainmission antennas by the flexing of deicer bootswhich are pneumatically actuated. Bleed air fromthe engine compressor is used to inflate the deicerboots and to supply vacuum, through the ejector sys-tem, for boot hold down during flight. A pressureregulator protects the system from over inflation.When the system is not in operation, a distributorvalve applies vacuum to the boots for hold down. Aselector switch allows automatic single cycle opera-tion or manual operation. To assure operation of thesystem in the event of failure of one engine, a checkvalve is incorporated in the bleed air line from eachengine to prevent loss of pressure through the com-pressor of the inoperative engine.

Wing ice lights allow the crew to detect ice for-mations. Ice protection of the engine is provided byinertial separation. Automatically cycled electro-thermal anti-icing boots are installed on the propel-ler blades. The engine air inlet leading edge lip isanti-iced by engine exhaust bleed. The fresh airinlets are located in sheltered areas and require nodeice protection.

2-41

TM 55-1510-221-10

Operation of the surface deice system inambient temperatures below -40°C cancause permanent damage to the deiceboots.

b. Operation.

(1.) Deice boots are intended to removeice after it has formed rather than prevent its forma-tion. For the most effective deicing operation, allowat least l/2 inch of ice to form on the boots beforeattempting ice removal. Very thin ice may crack andcling to the boots instead of shedding.

NOTE

Never cycle the system rapidly, this maycause the ice to accumulate outside thecontour of the inflated boots and preventice removal.

(2.) A three position switch on the over-head control panel placarded SURF DEICE MAN-UAL - OFF - SINGLE CYCLE AUTO, controls thedeicing operation. The switch is spring loaded toreturn to the OFF position from SINGLE CYCLEAUTO or MANUAL. When the SINGLE CYCLEAUTO position is selected, the distributor valveopens to inflate the wing boots. After an inflationperiod of approximately 6 seconds, an electronictimer switches the distributor to deflate the wingboots and a 4 second inflation begins in the horizon-tal stabilizer boots. When these boots have inflatedand deflated, the cycle is complete.

(3.) If the switch is held in the MANUALposition, the boots will inflate simultaneously andremain inflated until the switch is released. Theswitch will return to the OFF position whenreleased. After the cycle, the boots will remain in thevacuum hold down condition until again actuatedby the switch.

(4.) Either engine is capable of providingsufficient bleed air for all requirements of the sur-face deicer system. Check valves in the bleed air andvacuum lines prevent backflow through the systemduring single-engine operation. Regulated pressure isindicated on a gage, placarded PNEUMATIC PRES-SURE, located on the copilots subpanel.

2-52. ANTENNA DEICING SYSTEM.

a. Description. The antenna deice systemremoves ice accumulation from the dipole mission

2-42

antennas. The system consists of two ejector distrib-utor valves, a regulator, manifold, and flexible tub-ing. Control is accomplished through a timing cir-cuit and an antenna deice switch located on theoverhead control panel (fig. 2-12). Erosion resistanttape is applied to the surface of mission blade anten-nas not having deice boots.

b. Antenna Deice System Switch. The antennadeice system is controlled by a switch placardedANT DEICE, SINGLE - OFF - MANUAL locatedon the overhead control panel (fig. 2-12). The switchis spring loaded to return to the OFF position fromthe SINGLE or MANUAL position. When theswitch is set to the single position, the system willrun through one timed inflation-deflation cycle.When the switch is held in the MANUAL positionthe boots will inflate and remain inflated until theswitch is released.

c. Forward Wide Band Data Link AntennaRadome Anti-Ice. The forward wide band data linkantenna radome anti-ice system utilizes engine bleedair to prevent the formation of ice on the radome.The system is controlled by a switch placardedRADOME located on the overhead control panel.The circuit is protected by a circuit breaker plac-arded RADOME, located on the overhead circuitbreaker panel (fig. 2-26).

d. Operation.

(1.) Deice boots are intended to removeice after it has formed rather than prevent its forma-tion. For the most effective deicing operation, allowat least l/8 to l/4 inch of ice to form on the bootsbefore attempting ice removal. Very thin ice maycrack and cling to the boots instead of shedding.

NOTE

Never cycle the system rapidly, this maycause the ice to accumulate outside thecontour of the inflated boots and preventice removal.

2-53. PROPELLER ELECTROTHERMAL ANTI-ICESYSTEM.

a. Description. Electrothermal anti-ice bootsare cemented to each propeller blade to prevent iceformation or to remove the ice from the propellers.Each thermal boot consists of one outboard and oneinboard heating element, and receives electricalpower from the deice timer. This timer sends cur-rent to all propeller deice boots and prevents theboots from overheating by limiting the time eachelement is energized. Four intervals of approxi-mately 30 seconds each complete one cycle. Current

TM 55-1510-221-10

consumption is monitored by a propeller ammeteron the copilot’s subpanel. Two 20-ampere circuitbreakers placarded PROP ANTI-ICE LEFT andRIGHT and 5-ampere propeller control circuitbreaker placarded CONTR on the overhead circuitbreaker panel (fig. 2-26), protect the propeller elec-trothermal deice system during manual operation. A25 ampere circuit breaker placarded PROP AUTO,protects the system in automatic operation.

b. Automatic Operation. A control switch onthe overhead control panel placarded PROP - OFF- AUTO is provided to activate the automatic sys-tem. A deice ammeter above the pedestal registersthe amount of current (14 to 18 amperes) passingthrough the system being used. During AUTO oper-ation, power to the timer will be cut off if the cur-rent rises above 25 amperes. Current flows from thetimer to the brush assembly and then to the sliprings installed on the spinner backing plate. The sliprings carry the current to the deice boots on the pro-peller blades. Heat from the boots reduces the gripof the ice which is then thrown off by centrifugalforce, aided by the air blast over the propeller sur-faces. Power to the two heating elements on eachblade, the inner and outer element, is cycled by thetimer in the following sequence: right propeller outerelement, right propeller inner element, left propellerouter element, left propeller inner element. Loss ofone heating element circuit on one side does notmean that the entire system must be turned off.Proper operation can be checked by noting the cor-rect level of current usage on the ammeter. An inter-mittent flicker of the needle approximately each 30seconds indicates switching to the next group ofheating elements by the timer.

c. Manual Operation. The manual propellerdeice system is provided as a backup to the auto-matic system. A control switch located on the over-head control panel, placarded PROP - INNER -OUTER, controls the manual override relays. Whenthe switch is in the OUTER position, the automatictimer is overridden and power is supplied to theouter heating elements of both propellers simulta-neously. The switch is of the momentary type andmust be held in position until the ice has been dis-lodged from the propeller surface. After deicing withthe outer elements, the switch is to be held in theINNER position to perform the same function forthe inner elements of both propellers. The load-meters will indicate approximately a 5% increase ofload per meter when manual propeller deice is oper-ating. The propeller deice ammeter will not indicateany load in the manual mode of operation.

2-54. PITOT AND STALL WARNING HEAT SYS-TEM.

Pitot heat should not be used for morethan 15 minutes while the aircraft is onthe ground. Overheating may damage theheating elements.

a. Pitot Heat. Heating elements are installedin both pitot masts, located on the nose. Each heat-ing element is controlled by an individual switchplacarded PITOT - ON - LEFT or RIGHT, locatedon the overhead control panel (fig. 2-12). It is notadvisable to operate the pitot heat system on theground except for testing or for short intervals oftime to remove ice or snow from the mast. Circuitprotection is provided by two 7 l/2 ampere circuitbreakers, placarded PITOT HEAT, on the overheadcircuit breaker panel (fig. 2-26). The “true airspeedtemp probe“ heat control circuit is also protected bythis circuit breaker. If either left or right pitot heatis on, the“ true airspeed temp probe“ heat will beon.

NOTE

The “TRUE AIRSPEED TEMP PROBE“is connected to the autopilot air datacomputer.

The heating elements protect the stallwarning lift transducer vane and faceplate from ice, however, a buildup of iceon the wing may change or disrupt theairflow and prevent the system from accu-rately indicating an imminent stall.

b. Stall Warning Heat. The lift transducer isequipped with anti-icing capability on both themounting plate and the vane. The heat is controlledby a switch located on the overhead control panelplacarded STALL WARN. The level of heat is mini-mal for ground operation but is automaticallyincreased for flight operation through the landinggear safety switch. Circuit protection is provided bya 15-ampere circuit breaker, placarded STALLWARN, on the overhead circuit breaker panel (fig.2-26).

2-43

TM 55-1510-221-10

2-55. STALL WARNING SYSTEM.

The stall warning system consists of a trans-ducer, a lift computer, a warning horn, and a testswitch. Angle of attack is sensed by aerodynamicpressure on the lift transducer vane located on theleft wing leading edge. When a stall is imminent, theoutput of the transducer activates a stall warninghorn. The system has preflight test capabilitythrough the use of a switch placarded STALLWARN TEST - OFF - LDG GEAR WARN TESTon the right subpanel. Holding this switch in theSTALL WARN TEST position actuates the warninghorn by moving the transducer vane. The circuit isprotected by a 5-ampere circuit breaker, placardedSTALL WARN, on the overhead circuit breakerpanel.

2-56. BRAKE DEICE SYSTEM.

a. Description. A heated-air brake deice sys-tem may be used in flight with gear retracted orextended, or on the ground. When activated, hot airis diffused by means of a manifold assembly overthe brake discs on each wheel. Manual and auto-matic controls are provided. There are two primaryoccasions which require brake deicing. The first iswhen an aircraft has been parked in a freezing atmo-sphere allowing the brake systems to become con-taminated by freezing rain, snow, or ice, and the air-craft must be moved or taxied. The second occasionis during flight through icing conditions with wetbrake assemblies presumed to be frozen, which mustbe thawed prior to landing to avoid possible tiredamage and loss of directional control. Hot air forthe brake deice system comes from the compressorstage of both engines obtained by means of a sole-noid valve attached to the bleed air system whichserves both the surface deice system and the pneu-matic systems operation.

b. Operation. A switch on the overhead con-trol panel, placarded BRAKE DEICE, controls thesolenoid valve by routing power through a controlmodule box under the aisleway floorboards. Whenthe switch is on, power from a 5-ampere circuitbreaker on the overhead circuit breaker panel isapplied to the control module. A lo-minute timerlimits operation and avoids excessive wheel welltemperatures when the landing gear is retracted. Thecontrol module also contains a circuit to the greenBRAKE DEICE ON annunciator light, and has aresetting circuit interlocked with the gear uplockswitch. When the system is activated, the BRAKEDElCE ON light should be monitored and the con-trol switch selected OFF after the light extinguishes- otherwise, on the next gear extension the systemwill restart without pilot action. The control switchshould also be selected OFF, if deice operation fails

2-44

to self-terminate after 10 minutes. If the automatictimer has terminated brake deicer operation afterthe last retraction of the landing gear, the landinggear must be extended in order to obtain furtheroperation of the system.

(1.) The L BL AIR FAIL or R BL AIRFAIL annunciator lights may momentarily illumi-nate during simultaneous operation of the surfacedeice and brake deice systems at low N1 speeds. Ifthe lights immediately extinguish, they may be disre-garded.

(2.) During certain ambient conditions,use of the brake deice system may reduce availableengine power, and during flight will result in a TGTrise of approximately 20°C. Appropriate perfor-mance charts should be consulted before brake deicesystem use. If specified power cannot be obtainedwithout exceeding limits, the brake deice systemmust be selected off until after takeoff is completed.TGT limitations must also be observed when settingclimb and cruise power. The brake deice system isnot to be operated above 15°C ambient temperature.The system is not to be operated for longer than 10minutes (one deicer cycle) with the landing gearretracted. If operation does not automatically termi-nate after approximately 10 minutes following gearretraction, the system must be manually selected off.During periods of simultaneous brake deice and sur-face deice operation, maintain 85% N1 or higher. Ifinadequate pneumatic pressure is developed forproper surface deicer boot inflation, select the brakedeice system off. Both sources of pneumatic bleedair must be in operation during brake deice systemuse. Select the brake deice system off during single-engine operation. Circuit protection is provided bya 5-ampere circuit breaker, placarded BRAKEDEICE, on the overhead circuit breaker panel (fig.2-26).

2-57. FUEL SYSTEM ANTI-ICING.

a. Description. An oil-to-fuel heat exchanger,located on each engine accessory case, operates con-tinuously and automatically to heat the fuel suffi-ciently to prevent freezing of any water in the fuel.No controls are involved. Two external fuel ventsare provided on each wing. One is recessed to pre-vent ice formation; the other is electrically heatedand is controlled by two toggle switches on the over-head control panel placarded FUEL VENT ON,LEFT and RIGHT (fig. 2-12). They are protected bytwo 5-ampere circuit breakers, placarded FUELVENT HEAT, RIGHT or LEFT, located on theoverhead circuit breaker panel (fig. 2-26). Each fuelcontrol unit is protected against ice. The pneumaticgovernor for each fuel control unit is electricallyheated, and protected by two 7 1/2-ampere circuit

breakers located on the overhead circuit breakerpanel placarded FUEL CONTR HEAT, LEFT orRIGHT (fig. 2-26).

To prevent overheat damage to electri-cally heated anti-ice jackets, the FUELVENT heat switches should not be turnedON unless cooling air will soon pass overthe jackets.

b. Normal Operation. For normal operation,switches for the FUEL VENTS anti-ice circuits areturned ON as required during the BEFORE TAKE-OFF procedures (Chapter 8).

2-58. WINDSHIELD ELECTROTHERMAL ANTI-ICE SYSTEM.

a. Description. Both pilot and copilot wind-shields are provided with an electrothermal anti-icesystem. Each windshield is part of an independentelectrothermal anti-ice system. Each system is com-prised of the windshield assembly with heating wiressandwiched between glass panels, a temperature sen-sor attached to the glass, an electrothermal control-ler, two relays, a control switch, and two circuitbreakers. Two switches, placarded WSHLD ANTI-ICE NORMAL - OFF - HI - PILOT, COPILOT,located on the overhead control panel (fig. 2-l 2)control system operation. Each switch controls oneelectrothermal windshield system. The circuits ofeach system are protected by a 5-ampere circuitbreaker and a 50-ampere circuit breaker which arenot accessible to the flight crew. The 50-ampere cir-cuit breakers are located in the power distributionpanel under the floor ahead of the main spar. The5-ampere circuit breakers are located on panels for-ward of the instrument panel.

b. Normal Operation. Two levels of heat areprovided through the three position switches plac-arded NORMAL in the aft position, OFF in the cen-ter position, and HI after lifting the switch over adetent and moving it to the forward position. In theNORMAL position, heat is provided for the majorportion of each windshield. In the HI position, heatis provided at a higher watt density to a smaller por-tion of the windshield. The lever lock feature pre-vents inadvertent switching to the HI position dur-ing system shutdown.

2-59. PRESSURIZATION SYSTEM.

a. Description. A mixture of bleed air fromthe engines, and ambient air, is available for pres-

TM 55-1510-221-10

surization to the cabin at a rate of approximately 10to 17 pounds per minute. The flow control unit ofeach engine controls the bleed air from the engine tomake it usable for pressurization by mixing ambientair with the bleed air depending upon aircraft alti-tude and ambient temperature. On takeoff, excessivepressure bumps are prevented by landing gear safetyswitch actuated solenoids incorporated in the flowcontrol units. These solenoids, through a time delay,stage the input of ambient air flow by allowingambient air flow introduction through the left flowcontrol unit first, ten seconds later, air flow throughthe right flow control unit. The bleed air switches,located on the overhead control panel (fig. 2-12)operate an integral electric solenoid which controlsthe bleed air to the firewall shutoff valves.

b. Pressure Differential. The pressure vessel isdesigned for a normal working pressure differentialof 6.0 PSI, which will provide a cabin pressure alti-tude of 3870 feet at an aircraft altitude of 20,000feet, and a nominal cabin altitude of 9840 feet at anaircraft altitude of 31,000 feet.

c. Cabin Altitude and Rate-of-Climb Control-ler. A control panel is installed on the copilot’s sideof the subpanel (fig, 2-6) for operation of the system.A knob, placarded INC RATE controls the rate ofchange of pressurization. A control, placardedCABIN CONTROLLER is used to set the desiredcabin altitude. For proper cabin pressurization, theCABIN CONTROLLER should be set 500 feetabove cruise altitude. For landing select 500 feetabove field pressure altitude. The selected altitude isdisplayed on a mechanically coupled dial above thecontrol, placarded CABIN ALT-FT. Mechanicallycoupled to the cabin altitude dial, placardedACFTX 1000. This dial indicates the maximum alti-tude the aircraft may be flown at to maintain thedesired cabin altitude without exceeding the designpressure differential. A switch, placarded CABINPRESS DUMP-PRESS-TEST, is provided to controlpressurization. The switch is spring loaded to thePRESS position. In the DUMP position, the safetyvalve will be opened and the cabin will be depressu-rized to the aircraft altitude. In the PRESS position,cabin altitude is controlled by the CABIN CON-TROLLER control. In the TEST position, the land-ing gear safety switch is bypassed to enable testing ofthe pressurization system on the ground. Operatinginstructions are contained in Chapter 8.

d. Cabin Rate-of-Climb Indicator. An indica-tor, placarded CABIN CLIMB, is installed on thecopilot’s side of the instrument panel (fig. 2-29). Thecabin rate-of-climb controller is calibrated in thou-sands-of-feet per-minute change in cabin altitude.

e. Cabin Altitude Indicator. An indicator,placarded CABIN ALT, is installed in the instru-

2-45

TM 55-1510-221-10

ment panel (fig. 2-29) above the cabin rate-of-climb integral electric solenoid which controls the bleed airindicator. The longer needle indicates aircraft alti- to the firewall shutoff valves. A normally open sole-tude in thousands-of-feet on the outside dial. The noid operated by the landing gear safety switch con-shorter needle indicates pressure differential in PSI trols the introduction of ambient air flow to theon the inner dial. Maximum differential is 6.1 PSI. cabin on takeoff.

f. Outflow Valve. A pneumatically operatedoutflow valve, located on the aft pressure bulkhead,maintains the selected cabin altitude and rate-of-climb commanded by the cabin rate-of-climb andaltitude controller on the copilot’s instrument panel.As the aircraft climbs, the controller modulates theoutflow valve to maintain a selected cabin rate ofclimb and increases the cabin differential pressureuntil the maximum cabin pressure differential isreached. At a cabin altitude of 12,500 feet a pressureswitch mounted on the back of the overhead controlpanel completes a circuit to illuminate a red warningannunciator light, ALT WARN, to warn of opera-tion requiring oxygen. This light is protected by a5-ampere breaker, placarded PRESS CONTR.

(1.) The unit receives bleed air from theengine into an ejector which draws ambient air intothe nozzle of the venturi. The mixed air is thenforced into the bleed air line routed to the cabin.

(2.) Bleed air flow is controlled automati-cally. When the aircraft is on the ground, circuitryfrom the landing gear safety switch prevents ambi-ent air from entering the flow control unit to pro-vide maximum heating.

g. Pressurization Safety Valve. Before takeoff,the safety valve is open with equal pressure betweenthe cabin and the outside air. The safety valve closeson liftoff if the CABIN PRESS CONTR switch onthe instrument panel is in the PRESS mode. Thesafety valve adjacent to the outflow valve providespressure relief in the event of failure of the outflowvalve. This valve is also used as a dump valve andis opened by vacuum which is controlled by a sole-noid valve operated by the cabin pressure dumpswitch adjacent to the controller. It is also wiredthrough a landing gear safety switch. If either ofthese switches is open, or the vacuum source or elec-trical power is lost, the safety valve will close toatmosphere except at maximum differential pressureof 6.1 PSI. A negative pressure relief diaphragm isalso incorporated into the outflow and safety valvesto prevent outside atmospheric pressure fromexceeding cabin pressure during rapid descent.

(3.) The bleed air firewall shutoff valve inthe control unit is a spring loaded, bellows operatedvalve that is held in the open position by bleed airpressure. When the electric solenoid is shut off, orwhen bleed air diminishes on engine shutdown (inboth cases the pressure to the firewall shutoff valveis cut off), the firewall valve closes.

2-60. OXYGEN SYSTEM.

h. Drain. A drain in the outflow valve staticcontrol line is provided for removal of accumulatedmoisture. The drain is located behind the lower side-wall upholstery access panel in the baggage sectionof the aft compartment.

i. Flow Control Unit. A flow control unit for-ward of the firewall in each nacelle controls bleed airflow and the mixing of ambient air to make up thetotal air flow to the cabin for pressurization, heating,and ventilation. The bleed air switches located onthe overhead control panel (fig. 2-12) operates an

a. Description. The oxygen system (fig. 2-19)is provided primarily as an emergency system, how-ever, the system may be used to provide supplemen-tal (first aid) oxygen. Two 70 cubic foot capacityoxygen supply cylinders charged with aviator’sbreathing oxygen are installed in the unpressurizedportion of the aircraft behind the aft pressure bulk-head. The pilot and copilot positions are equippedwith diluter demand type regulators, which mix theproper amount of oxygen for a given amount of airat altitude. Also a first aid oxygen mask is providedin the cabin. Oxygen system pressure is shown bytwo gages placarded OXYGEN SUPPLY PRES-SURE, located aft of the pilot’s oxygen regulatorcontrol panel. Two pressure reducers, located in theunpressurized portion of the aircraft behind the aftbulkhead, lower the pressure in the system to 400PSI, and route oxygen to the regulator control pan-els. Both cylinders are interconnected, so refillingcan be accomplished through a single tiller valvelocated on the aft right side of the fuselage exterior.A pressure gage is mounted in conjunction with thefiller valve, and each cylinder has a pressure gage.Table 2-4 shows oxygen flow planning rates vs alti-tude. Table 2-5 shows oxygen duration capacities ofthe system.

2-46

TM 55-1510-221-10

--------------LOW PRESSURE SYSTEM

H I G H P R E S S U R E S Y S T E MAPOO6599

Figure 2- 19. Oxygen System Schematic

2-47

TM 55-1510-221-10

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

CABIN PRESSURE CREW MASK CREW MASK PASSENGERALTITUDE IN FEET NORMAL 100% MASK

(DILUTER (1)DEMAND)

(1)

3 1,000 -0-( 2) 4.2 3.7 (3)

30,000 -0-( 2) 4.4 3.7 (3)29,000 -0-( 2) 4.7 3.7 (3)28,000 -0-( 2) 5.0 3.7 (3)27,000 -0-( 2) 5.3 3.7 (3)26,000 -0-( 2) 5.6 3.7 (3)25,000 -0-( 2) 5.9 3.724,000 -0-( 2) 6.2 3.7

23,000 -0-( 2) 6.6 3.722,000 -0-( 2) 6.9 3.72 1 ,000 -0-( 2) 7.2 3.720,000 3.6 7.6 3.719,000 3.9 7.9 3.718,000 4.2 8.3 3.717,000 4.5 8.7 3.7

16,000 4.8 9.1 3.715,000 5.1 9.5 3.714,000 5.4 10.0 3.713,000 5.8 10.4 3.712,000 6.1 10.9 3.711,000 6.5 11.3 3.7

10,000 6.9 11.9 3.7

NOTES:(1) Based on minute volume of 20 LPM-BTPS (Body Temperature and Pressure Saturated).(2) Use 100% oxygen at or above 20,000 feet.(3) Not recommended for other than emergency descent use above 25,000 feet.

If average climb or descent flows are desired, add the values between altitudes and divide by the numberof values used.

For example, to determine the average rate for a uniform descent between 25,000 feet and 15,000 feetperform the following.

(5.9 + 6.2 + 6.6 + 6.9 + 7.2 + 7.6 + 3.9 + 4.2 + 4.5 + 4.8 + 5.1) ÷ 11 = 5.7 LPM

This method is preferred over averaging the extremes an some flow characteristics vary in such a way asto yield as incorrect answer.

2-48

TM 55-1510-221-10

Table 2-5. Oxygen Duration In Minutes 140 Cubic Foot System

TWO MAN 3 1,000 100% 8.4CREW 25,000 100% 11.8

20,000 100% 15.220,000 NORMAL 7.415,000 100% 19.015,000 NORMAL 10.210,000 100% 23.8

CABIN CREW TOTALPRESSURE MASK FLOWALTITUDE CONDITION LPM-NTPD

TWO MANCREW PLUSONE PASS

10,0003 1,00025,00020,00020,00015,000

15,000

10,00010,000

(1) For 100% capacity of useable oxygen, 3,226 L.

NORMAL100%100%100%

NORMAL

100%NORMAL

100%NORMAL

13.8 233.712.1 266.615.5 208.118.9 170.010.9 295.922.7 142.1

13.9 232.1

27.5 117.317.5 184.3

DURATIONIN

MINUTES (1)

384.0273.3212.2448.0169.7316.2

135.5

2-61. OXYGEN DURATION EXAMPLE PROBLEM

WANTED

Duration in minutes of oxygen at 100% capac-ity.

KNOWN

Two man crew plus one passenger, cabin pres-sure altitude = 15,000 feet, crew masks, normal,100% capacity.

METHOD

Find “two man crew plus one pass” line, moveright then down to 15,000 - “normal” read “232.1”minutes.

WANTED

Duration of oxygen for previous example data at84% of capacity.

KNOWN

232.1 minutes duration at l00%, 84% capacity,total aircraft flow = 13.9 LPM.

METHOD

Multiply 232.1 X 0.84 = 194.9 minutes. or Mul-tiply 3,226 X 0.84 = 2709.8, divide by 13.9 LPM =194.9 minutes.

WANTED

Duration of oxygen for complement at othercabin pressure altitude, at less than 100% capacity.

KNOWN

Cylinder at 84% capacity, 100% capacity =3,226 L, cabin pressure altitude = 21,000 feet.

1 crew mask = 7.2 LPM ( 100%) 1 passengermask = 3.7 LPM

METHOD

Multiply 3,226 L X 0.84 = 2.709.8 L, multiply2 crew X 7.2 LPM = 14.4 LPM, multiply 1 passen-ger X 3.7 LPM, add 14.4 LPM crew plus 3.7 LPMpassenger = 18.1 LPM. Divide 3,226 L by 18.1LPM = 178.2 minutes.

2-62. OXYGEN CYLINDER CAPACITY EXAMPLEPROBLEM

WANTED

a. Percent of capacity at known pressure andtemperature.

b. Pressure when temperature decreases.

2-49

TM 55-1510-221-10

STABILIZED CYLINDER TEMPERATURE

500 1,000 1,500 2,000CYLINDER PRESSURE (PSIG)

Figure 2-20. Cylinder Capacity vs Pressure and Temperature

KNOWN

Pressure = 1,600 PSIG stabilized cylinder tem-perature is estimated at 20º C decreased stabilizedcylinder temperature is estimated at -30º C.

METHOD

a. Enter 1600 PSIG move up to 20º C line,move right to 84%.

b. Move left on 84% line to -30º C line, movedown to 1250 PSIG.

WANTED

100% capacity pressure at known temperature.

KNOWN

Temperature = -30º C.

METHOD

Move left along 100% line to -30º C line andmove down to 1420 PSIG.

(1.) Regulator control panels. Each regula-tor control panel contains a blinker-type flow indica-tor, a 500 PSI pressure gage, a red emergency pres-sure control lever placarded EMERGENCY -NORMAL - TEST MASK, a white diluter controllever placarded 100% OXYGEN - NORMAL OXY-

AP010340

GEN, and a green supply control lever placardedON - OFF. The diluter control lever selects eithernormal or 100% oxygen, but acts to select only whenthe emergency pressure control lever is in the NOR-MAL position.

When not in use, the diluter control levershould be left in the 100% OXYGENposition to prevent regulator contamina-tion.

(2.) The emergency pressure control leverhas three positions. Two positions control oxygenconsumption for the individual using oxygen, andthe remaining position serves for testing hose andmask integrity. In the EMERGENCY position, thecontrol lever causes 100% oxygen to be delivered ata safe, positive pressure. In the NORMAL position,the lever allows delivery of normal or 100% oxygen,depending upon the selection of the diluter controllever. In TEST MASK position, 100% oxygen atpositive pressure is delivered to check hose andmask integrity.

(3.) The 500 PSI oxygen pressure gageprovided on the oxygen control panels should never

2-50

TM 55-1510-221-10

indicate over 400 PSI. If the pressure exceeds 400 vided with a cord for connecting with the micro-PSI, a malfunction of the pressure reducer is indi- phone jack. To test mask and hose integrity, thecated. Whenever oxygen is inhaled, a blinker-vane individual places the supply control lever on the reg-slides into view within the flow indicator window, ulator control panel to the ON position, puts on andshowing that oxygen is being released. When oxygen adjusts his mask, selects TEST MASK position, andis exhaled, the blinker vane vanishes from view. checks for leaks.

NOTE

Check to insure that the OXYGEN SUP-PLY PRESSURE gage registers adequatepressure before each flight. When oxygenis in use, a check of the supply pressureshould be made at intervals during flightto note the quantity available and toapproximate the supply duration. Theoutside temperature is reduced as an air-craft ascends to higher altitudes. Oxygencylinders thus cooled by temperaturechange will show a pressure drop. Thistype of drop in pressure will raise againupon return to a lower or warmer alti-tude. A valid cause for alarm would bethe rapid loss of oxygen pressure when theaircraft is in level flight or descending;should this condition arise, descend asrapidly as possible to altitude which doesnot require the use of oxygen.

W A R N I N G

Pure oxygen will support combustion. Donot smoke while oxygen is in use.

W A R N I N G

If any symptoms occur suggestive of theonset of hypoxia, immediately set theemergency pressure control lever to theEMERGENCY position and descendbelow 10,000 feet. Whenever carbonmonoxide or other noxious gas is presentor suspected, set the dilutor control leverto 100% OXYGEN and continue breath-ing undiluted oxygen until the danger ispast.

e. Oxygen masks. Oxygen masks for the pilotand copilot are provided as personal equipment. Toconnect a mask into the oxygen system, the individ-ual connects the line attached to the mask to theflexible hose which is attached to the cockpit side-wall. The microphone in the oxygen mask is pro-

f. Normal Operation. Oxygen pressure ismaintained at all times to the regulator control pan-els if the cylinder shut-off valves are on and if thereis pressure in the cylinders. Each individual placesthe supply lever (green) on his regulator controlpanel to the ON position, and the diluter lever(white) to the NORMAL OXYGEN position.

g. Emergency Operation. For emergency oper-ation, the affected crew member selects the EMER-GENCY position of the emergency pressure controllever (red) on his regulator control panel. This selec-tion provides 100% oxygen at a positive pressure,regardless of the position of the diluter control leveron his panel.

h. First Aid Operation. A first aid oxygenmask is installed in the aft cabin area as a supple-mental or emergency source of oxygen. The mask isstowed behind an overhead cover placarded FIRSTAID OXYGEN -PULL. Removing the cover allowsthe mask to drop out of the container, exposing amanual control valve, which releases oxygen to themask when placed in the ON position. After usingthe mask, the manual valve in the container must beturned OFF before stowing the mask and replacingthe cover.

2-63. WINDSHIELD WIPERS.

a. Description. Two electrically operatedwindshield wipers, are provided for use at takeoff,cruise and landing speed. A rotary switch (fig. 2-12)placarded WINDSHIELD WIPER, located on theoverhead control panel, selects mode of windshieldwiper operation. An information placard above theswitch states: DO NOT OPERATE ON DRYGLASS. Function positions on the switch, as readclockwise, are placarded: PARK - OFF - SLOW -FAST. When the switch is held in the spring-loadedPARK setting, the blades will return to their normalinoperative position on the glass, then, whenreleased, the switch will return to OFF position ter-minating windshield wiper operation. The FASTand SLOW switch positions are separate operatingspeed settings for wiper operation. The windshieldwiper circuit is protected by one lo-ampere circuitbreaker, placarded WSHLD WIPER, located on theoverhead circuit breaker panel (fig. 2-26).

2-51

TM 55-1510-221-10

Do not operate windshield wipers on dryglass. Such action can damage the linkageas well as scratch the windshield glass.

b. Normal Operation. To start, turn WIND-SHIELD WIPER switch to FAST or SLOW speed,as desired. To stop, turn the switch to the PARKposition and release. The blades will return to theirnormal inoperative position and stop. Turning theswitch only to the OFF position will stop the wind-shield wipers, without returning them to the normalinactive position.

2-64. FERRY CHAIR.

For ferry purposes, a forward facing chair witha lap belt is attached to floor tracks at fuselage sta-tion 211.87 (fig. 2-2).

2-65. CIGARETTE LIGHTERS AND ASH TRAYS.

The pilot and copilot have individual cigarettelighters and ash trays mounted in escutcheons out-board of their seats. The cigarette lighters are pro-tected by a 5-ampere circuit breaker, placardedCIGAR LIGHTER, on the overhead circuit breakerpanel (fig. 2-26).

2-66. CHEMICAL TOILET.

a. Description. A side-facing chemical toilet(figure 2-2) is installed in the aft cabin area. Twohinged lid half-sections must be raised to gain accessto the toilet. Waste is stored within a removable

2-69.

a.

Section VIII. HEATING, VENTILATION, COOLING, AND ENVIRONMENTAL Control System

HEATING SYSTEM. ducted to the cockpit outlets, windshield defroster

Description. Warm air for heating the cock-pit and mission avionics compartments and warmwindshield defrosting air is provided by bleed air

container located below the seat in the cabinetassembly. This non-flushing system uses a dry chem-ical preparation to deodorize the stored waste. A toi-let tissue dispenser is contained in a slide-out com-partment on the forward side of the toilet cabinet. Abox of disposable waste container liners and a boxof chemicial deodorant packets are also stored in thecabinet.

b. Operation. During use, a removable, throw-away plastic liner is attached to the waste container.After use, dry chemical deodorant obtained from thestorage cabinet is deposited on the waste and thehinged lid sections are closed over the cavity. Aftereach flight, the waste container must be removed,emptied, relined, replaced in the cabinet and othertoilet items are resupplied as needed,

2-67. SUN VISORS.

When adjusting the sun visors, grasp onlyby the top metal attachment to avoiddamage to the plastic shield.

A sun visor is provided for the pilot and copilotrespectively (fig. 2-8). Each visor is manually adjust-able. When not needed as a sun shield, each visormay be manually rotated to a position flush with thetop of the cockpit so that it does not obstruct viewthrough the windows.

2-68. RELIEF TUBE.

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

outlets, and to the floor outlets in the mission avion-ics compartment. The environmental system isshown in figure 2-21. placarded WSHLD W1

from both engines. Engine bleed air is combinedwith ambient air in the heating and pressurizationflow control unit in each nacelle. If the mixed bleedair is too warm for cockpit comfort, it is cooled bybeing routed through an air-to-air heat exchangerlocated in the forward portion of each inboard wing.If the mixed bleed air is not too warm, the air-to-airheat exchangers are bypassed. The mixed bleed airis then ducted to a mixing plenum, where it is mixedwith cabin recirculated air. The warm air is then

(1.) Bleed airflow control unit. A bleed airflow control unit, located forward of the firewall ineach engine nacelle controls the flow of bleed airand the mixing of ambient air to make up the totalairflow to the cabin for heating, windshield defrost-ing, pressurization and ventilation. The unit is fullypneumatic except for an integral electric solenoidfirewall shutoff valve, controlled by the bleed airswitches located on the overhead control panel (fig.

2-52

TM 55-1510-221-10

Figure 2-2 1. Environmental System

2-53

TM 55-1510-221-10

2-l 2) and a normally open solenoid valve operatedby the right landing gear safety switch.

(2.) Pneumatic bleed air shutoff valve. Apneumatic shutoff valve is provided in each nacelleto control the flow of bleed air to the surface,antenna and brake deice systems. These valves arecontrolled by the bleed air valve switches located onthe overhead control panel (fig. 2-12).

(3.) Bleed air valve switches. The bleed airflow control unit shutoff valve and pneumatic bleedair shutoff valves are controlled by two switchesplacarded BLEED AIR VALVE - OPEN - ENVIROOFF - PNEU & ENVIRO OFF, located on the over-head control panel (fig. 2-12). When set to the openposition, both the environmental flow control unitshutoff valve and the pneumatic shutoff valve areopen; when set to the ENVIRO OFF position, theenvironmental flow control unit shutoff valve isclosed, and the pneumatic bleed air valve is open; inthe PNEU & ENVIRO OFF position, both areclosed. For maximum cooling on the ground, turnthe bleed air valve switches to the ENVIRO OFFposition.

(4.) Cabin temperature mode selectorswitch. A switch placarded CABIN TEMP MODE -MAN COOL - MAN HEAT -OFF - AUTO - A/CCOLD OPN - -25°C to 10°C located on the over-head control panel, controls cockpit and missionavionics compartment heating and air conditioning.When the cabin temperature mode selector switch isset to the AUTO position, the heating and air condi-tioning systems are automatically controlled. Con-trol signals from the temperature control box aretransmitted to the bleed air heat exchanger bypassvalves. Here the temperature of the air flowing tothe cabin is regulated by the bypass valves control-ling the amount of air bypassing the heat exchang-ers. When the temperature of the cabin has reachedthe temperature setting of the cabin temperaturecontrol rheostat, the automatic temperature controlallows hot air to bypass the air-to-air exchangersadmitting hot air into the cabin. When the bypassvalves are in the fully closed position, allowing noair to bypass the heat exchangers, the air conditionerbegins to operate, providing additional cooling.When the cabin temperature mode selector switch isset to the A/C COLD OPN position, the air condi-tioning system is in continuous operation. The cabintemperature control rheostat, in conjunction withthe cabin temperature control sensor, provides regu-lation of cockpit and mission equipment compart-ment temperature. Bleed air heat is added asrequired to maintain the temperature selected by thecabin temperature control rheostat.

(5.) Cabin temperature control rheostat. Acontrol knob placarded CABIN TEMP - INCR,

2-54

located on the overhead control panel (fig. 2-12),provides regulation of cabin temperature when thecabin temperature mode selector switch is set to theAUTO or the A/C COLD OPN position. A tempera-ture sensing unit in the cabin, in conjunction withthe setting of the cabin temperature control rheostat,initiates a heat or cool command to the temperaturecontroller for the desired cockpit or mission avionicscompartment environment.

(6.) Manual temperature control switch. Aswitch placarded MANUAL TEMP - INCR -DECR, located on the overhead control panel (fig.2-12), controls cockpit and mission avionics com-partment temperature with the cabin temperaturemode selector switch set to the MAN HEAT posi-tions. The manual temperature control switch con-trols cockpit and mission avionics temperature byproviding a means of manually changing the amountthat the bleed air bypass valves are opened orclosed. To increase cabin temperature the switch isheld to the INCR position. To decrease cabin tem-perature, the switch is held to the DECR position.Approximately 30 seconds per valve is required todrive the bypass valves to the fully open or fullyclosed position. Only one valve moves at a time.

(7.) Forward vent blower switch. The for-ward vent blower is controlled by a switch placardedVENT BLOWER - AUTO - LO - HI, located on theoverhead control panel (fig. 2-12). In the auto posi-tion the fan will run at low speed except when thecabin temperature mode selector switch is set to theOFF position, in this case the blower will not oper-ate.

(8.) Aft vent blower switch. The aft ventblower is controlled by a switch placarded AFTVENT BLOWER - OFF - AUTO - ON, located onthe overhead control panel (fig. 2-l 2). The singlespeed blower operates automatically through thecabin temperature mode selector switch when the aftvent blower switch is placed in the AUTO position.The blower runs continuously when the switch isplaced in the ON position, In the OFF position, theblower will not operate.

b. Automatic Heating Mode.

1. Bleed air valve switches - OPEN, LEFT andRIGHT.

2. Cabin temperature mode selector switch -AUTO.

3. Cabin temperature control rheostat - Asrequired.

4. Cabin, cockpit and defrost air knobs - Asrequired

c.

1.

2.

3.

4.

5.

2-70. AIR CONDITIONING SYSTEM.

Manual heating mode.

Bleed air valve switches - OPEN, LEFT andRIGHT.

Cabin temperature mode selector switch -MAN HEAT.

Vent blower switches - As required.

Manual temperature switch - As required.

Cabin, cockpit and defrost air knobs - Asrequired.

a. Description. Cabin air conditioning is pro-vided by a refrigerant gas vapor cycle refrigerationsystem consisting of a belt driven, engine mountedcompressor, installed on the No.2 engine accessorypad, refrigerant plumbing, N1 speed switch, high andlow pressure protection switches, condenser coil,condenser-under-pressure switch, condenser blower,forward and aft evaporator, receiver-dryer, expan-sion valve and a bypass valve. The plumbing fromthe compressor is routed through the right inboardwing leading edge to the fuselage and then forwardto the condenser coil, receiver-dryer, expansionvalve, bypass valve, and forward evaporator, whichare located in the nose of the aircraft. A 7 1/2-ampere circuit breaker placarded AIR CONDCONTR, located on the overhead control panel (fig.2-12), protects the compressor clutch circuit.

(1.) Forward evaporator. The forwardevaporator and blower supplies the cockpit, forwardceiling outlets, and forward floor outlets. The for-ward evaporator blower has a high speed which canbe selected by setting the VENT BLOWER switch,located on the overhead control panel (fig. 2-12), tothe HI position. The forward vent blower is pro-tected by a circuit breaker on the DC power distri-bution panel, located below the aisleway floor for-ward of the main spar.

(2.) Aft evaporator. The aft evaporator andblower are located in the fuselage center aisle equip-ment bay aft of the rear spar. Environmental air iscirculated through the evaporator in either manualor automatic control mode. The rear evaporatorsupplies the aft ceiling outlets, rear floor outlets, andtoilet compartment. Rear evaporator blower is pro-tected by a circuit breaker on the DC power distri-bution panel located below the aisleway floor for-ward of the main spar.

(3.) High and low pressure limit switches.High and low pressure limit switches are provided toprevent compressor operation beyond operationallimits. When the low or high pressure switches areactivated, compressor operation will be terminated.

TM 55-1510-221-10

When compressor operation has been terminated bylimit switch activation, the system should be thor-oughly checked before returning it to service.

(4.) Thermal sense switch. A thermal senseswitch is installed on the forward evaporator. Thissense switch actuates a hot gas bypass valve whichbypasses a portion of the refrigerant from the for-ward evaporator, thereby preventing icing of theevaporator.

(5.) Condenser blower. A vane-axialblower draws air through the condenser on theground as well as in flight. The current limiter forthe blower is located on the DC distribution panelbelow the aisleway floor forward of the main spar.When the cabin temperature mode selector switch isset to the A/C COLD OPN position, the condenserblower will be off, and will remain off until the con-denser blower control high pressure switch senses acompressor discharge pressure equal to the pressureit is set to. The condenser blower will then remainin operation until the low pressure switch senses thatthe system pressure has dropped to the pressure it isset to.

(6.) Air conditioning cold operation bypassvalve. Selecting the A/C COLD OPN mode on thecabin temperature mode selector switch permits theoperation of the air conditioning system by overrid-ing the refrigerant low pressure switch. This allowsthe air conditioning system to operate in the manualmode. Starting the compressor in this optional modeat low ambient temperature will decrease the opera-tional life of the compressor by five hours each timethe air conditioning system is started using thismode (A/C COLD OPN). If the air conditioning sys-tem has been operating in the normal mode duringflight, and due to decreasing ambient temperaturesmake it necessary to switch to the A/C cold opera-tion mode, there will be no degradation in the meantime between failure for the compressor.

(7.) Air conditioner cold operation advisoryannunciator light. A green advisory annunciatorlight placarded A/C COLD OPN, located on thecaution/advisory annunciator panel (fig. 2.6), illumi-nates when the air conditioning system is operatingin cold mode, or when ambient temperatures requireswitching to cold mode if air conditioning systemoperation is to be continued.

b. Normal Operation.

(1.) Automatic cooling mode.

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

2. Cabin temperature mode selectorswitch - AUTO.

2-55

TM 55-1510-221-10

3. Cabin temperature control rheo-stat - As required.

4. Cabin, cockpit and defrost airknobs - As required.

(2.) Manual cooling mode.


NOTE

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

2.

(3.) Air conditioning cold operation mode.(Used if ambient-25º C).

Cabin temperature mode selectorswitch - MAN COOL.

temperature is between 10°C and

NOTE

Setting the cabin temperature mode selec-tor switch to the A/C COLD OPN posi-tion at ambient temperatures below -25°Cmay cause the air conditioning system toexceed the compressor low pressure limitswitch setting, terminating compressoroperation, and thereby rendering the sys-tem inoperative for the remainder of theflight.


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

3. Cabin temperature control rheo-stat - As required.

4. Cabin, cockpit and defrost airknobs - As required.

2-71. UNPRESSURIZED VENTILATION.

Ventilation is provided by two sources. Onesource is through the bleed air heating system inboth the pressurized and unpressurized mode. Thesecond source of ventilation is obtained from ramair through the condenser section in the nosethrough a check valve in the vent blower plenum.Ventilation from this source is in the unpressurizedmode only with the CABIN PRESS DUMP switchin the DUMP position. The check valve closes dur-ing pressurized operation. Ram air ventilation is dis-

tributed through the main ducting system to all out-lets. Ventilation air, ducted to each individual eye-ball cold air outlet, can be directionally controlledby moving the ball in the socket. Volume is regu-lated by twisting the outlet to open or close thevalve.

2-72. ENVIRONMENTAL CONTROLS.

An environmental control section on the over-head control panel (fig. 2-12) provides for automaticor manual control of the system. This section con-tains all the major controls of the environmentalfunction including bleed air valve switches, a ventblower control switch, an aft vent blower switch, amanual temperature switch for control of the heatexchanger valves, a cabin temperature level control,and the cabin temp mode selector switch for select-ing automatic heating or cooling or manual heatingor cooling. Four additional manual controls on themain instrument subpanels may be utilized for par-tial regulation of cockpit comfort when the cockpitpartition door is closed and the cabin comfort levelis satisfactory.

a. Heating Mode.

(1.) If the cockpit is too cold:

1. Pilot and copilot air knobs - Asrequired.

2. Defrost air knob - As required.

3. Cabin air knob - Pull out in smallincrements. Allow 3 -5 minutesafter each adjustment for systemto stabilize.

(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 asrequired.

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 - Outas required.

2-56

TM 55-1510-221-10

2. Cabin air knob. Close in smallincrements. Allow 3 - 5 minutesafter each adjustment for systemto stabilize. If CABIN AIR knobis completely closed beforeobtaining satisfactory cockpitcomfort, it may be necessary toplace the aft vent blower switch inthe ON position to activate theaft evaporator to recirculate cabinair.

c. Automatic Mode Control. When the AUTOmode is selected on the cabin temperature modeselector switch, the heating and air conditioning sys-tems are automatically controlled. When the temper-ature of the cabin has reached the selected setting,the automatic temperature control allows heated airto bypass the air-to-air exchangers in the wing centersection. By-pass air is hot. Heat exchanger air iscooled to approximately 30°F above ambient (out-side) air. The warm bleed air is mixed with thecooled air. The rear evaporator picks up recirculatedcabin air only.

NOTE

The automatic mode control works asdescribed, when the cabin is being cooledby bleed air. However, when the cabin isheated with bleed air and the selectedtemperature is reached, hot bleed airroutes through the heat exchanger forcooling in order to maintain the desiredtemperature.

(1.) When the automatic control drivesthe environmental system from a heat mode to acooling mode, the bypass valves close. When the leftbypass valve reaches a fully closed position, therefrigeration system will begin cooling, provided theright engine N1 speed is above 65%. When thebypass valve is opened to a position approximately30º from full open, the refrigeration system will turnOff.

(2.) The CABIN TEMP - INCR controlprovides regulation of the temperature level in theautomatic mode. A temperature sensing unit in thecabin, in conjunction with the control setting, initi-ates a heat or cool command to the temperaturecontroller for desired cockpit and cabin environ-ment.

d. Manual Mode Control. With the cabin tem-perature mode selector in the MAN HEAT or MAN

COOL position, regulation of the cabin temperatureis accomplished manually with the MANUALTEMP switch.

(1.) In the MAN HEAT mode, the auto-matic system is overridden and the system is con-trolled by opening and closing the bypass valves(two) with the MANUAL TEMP - INCR - DECRswitch. To increase cabin temperature, hold theswitch at the INCR position, to decrease cabin tem-perature, hold the switch in the DECR position.Allow approximately 30 seconds per valve to drivethe bypass valves to the fully open or fully closedposition. Only one valve moves at a time.

(2.) With the cabin temperature selectorswitch in the MAN COOL position, the automatictemperature control system is bypassed. When theleft bypass valve reaches a fully closed position, therefrigeration system will begin cooling, provided theright engine N1 speed is above 65%. When thebypass valve is opened to a position approximately30º from full open, the refrigeration system will turnoff. Hold the MANUAL TEMP switch in the DECRposition approximately one minute to fully close air-to-air heat exchanger bypass valves.

(3.) Bleed air entering the cabin is con-trolled by bleed air valve switches placarded BLEEDAIR VALVE - OPEN - ENVIRO OFF -PNEU &ENVIRO OFF. When the switch is in the OPENposition, the environmental flow control unit andthe pneumatic valve are open. When the switch is inthe ENVIRO OFF position, the environmental flowcontrol unit is closed and the pneumatic bleed airvalve is open; in the PNEU & ENVIRO OFF posi-tion, both are closed. For maximum cooling on theground, turn the bleed air valve switches to theENVIRO OFF position.

(4.) The forward vent blower is controlledby a switch placarded VENT BLOWER - AUTO -LOW - HI. The HI and LOW positions regulate theblower to two speeds of operation. IN the AUTOposition, the fan will run at low speed except whenthe CABIN TEMP mode selector switch is placed inthe OFF position. In the OFF position, the blowerwill not operate.

(5.) The aft vent blower is controlled by aswitch placarded AFT VENT BLOWER - OFF -AUTO - ON. The single speed blower operates auto-matically through the CABIN TEMP mode selectorwhen the AFT VENT BLOWER switch is placed inthe AUTO position. The blower runs continuouslywhen the switch is placed in the ON position. In theOFF position, the blower will not operate.

2-57

TM 55-1510-221-10

Section IX. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM

2-73. DESCRIPTION.

The aircraft employs both direct current (DC)and alternating current (AC) electrical power. TheDC electrical power supply (fig. 2-22) is the basicpower system energizing most aircraft circuits. Elec-trical power is used to start the engines, to power thelanding gear and flap motors, and to operate thestandby fuel pumps, ventilation blower, lights andelectronic equipment. AC power is obtained fromDC power through inverters. The single phase ACpower system is shown in figure 2-23, and the threephase AC power system is shown in figure 2-24. Thethree sources of DC power consist of one 20 cell 34-ampere/hour battery and two 400-ampere starter-generators. DC power may be applied to the aircraftthrough an external power receptacle on the under-side of the right wing leading edge just outboard ofthe engine nacelle (refer to Section XII for GPUrequirements). The starter-generators are controlledby generator control units. The output of each gener-ator passes through a cable to the respective genera-tor bus (fig. 2-22). Other buses distribute power toaircraft DC loads, and derive power from the gener-ator buses. The generators are paralleled to balancethe DC loads between the two units. When one ofthe generating systems is not on- line, and no faultexists, all aircraft DC requirements may be suppliedeither by the other on-line generating system, or byan external power source, but not by both. Most DCdistribution buses are connected to both generatorbuses but have isolation diodes to prevent powercrossfeed between the generating systems, when con-nection between the generator buses is lost. Thus,when either generator is lost because of a groundfault, the operating generator will supply power forall aircraft DC loads except those receiving powerfrom the inoperative generator’s bus which cannotbe crossfed. When a generator is not operating,reverse current and over-voltage protection is auto-matically provided. Two inverters operating fromDC power produce the required single-phase ACpower. Three phase AC electrical power for inertialnavigation system and mission avionics is suppliedby two DC powered mission inverters. AC powermay be applied through an external power receptaclelocated on the left nacelle. The mission power sys-tem is shown in figure 2-25.

engine-driven 28 volt, 400-ampere starter-generators. Controls and indicators associated withthe DC supply system are located on the overheadcontrol panel (fig. 2-12) and consist of a single bat-tery switch (BATT), two generator switches (No.1GEN and No.2 GEN), and two volt-loadmeters.

a. Battery Switch. A switch, placarded BATT(fig. 2-12) is located on the overhead control panelunder the MASTER SWITCH. The BATT switchcontrols DC battery power to the aircraft bus systemthrough the battery relay, and must be ON to allowexternal power to enter aircraft circuits. When theMASTER SWITCH is placed down, the BATTswitch is forced OFF.

NOTE

With battery or external power removedfrom the aircraft electrical system, due tofault, power cannot be restored to the sys-tem until the BATT switch is moved toOFF/RESET, then ON.

b. Generator Switches. Two switches (fig.2-12), placarded No. 1 GEN and No.2 GEN arelocated on the overhead control panel under theMASTER SWITCH. The toggle switches controlelectrical power from the designated generator toparalleling circuits and the bus distribution system.Switch positions are placarded RESET, ON andOFF. RESET is forward (spring-loaded back to ON),ON is center, and OFF is aft. When a generator isremoved from the aircraft electrical system, dueeither to fault or from placing the GEN switch in theOFF position, the affected unit cannot have its out-put 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.2-12) which extends above the battery and generatorswitches. The MASTER SWITCH gangbar is movedforward when a battery or generator switch is turnedon. When moved aft, the bar forces each switch tothe OFF position.

2-74. DC POWER SUPPLY.

One nickel-cadmium battery furnishes DCpower when the engines are not operating. This 24-volt, 34-ampere/hour battery, located in the rightwing center section, is accessible through a panel onthe top of the wing. DC power is produced by two

d. Volt-Loadmeters. Two meters (fig. 2-12),on the overhead control panel display voltage read-ings and show the rate of current usage from left andright generating systems. Each meter is equippedwith a spring-loaded pushbutton switch which whenmanually pressed will cause the meter to indicatemain bus voltage. Each meter normally shows theoutput amperage reading from the respective genera-tor, unless the pushbutton switch is pressed toobtain the bus voltage reading. Current consumption

2-58

TM 55-1510-221-10

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

2-59

TM 55-1510-221-10

HF RCVR#1 VHF#1 VOR#1 RMI

#1 AVIONICS BUS

PILOT AUDIOTRANSPONDERUHFTACANINS CONTROL

AFCS DIRECTAP PWRVHF/AM/FM

#2 AVIONICS BUS

#2 VOR COPILOT AUDIOADF RADAR

#2 RMI RADAR-NAV

SERVO DCRADIO RELAYCOPILOT ALTBU VOW

#1 DUAL FED BUS

ANN IND#1 CHIP DETR#1 QTY IND#1 QTY WARN#1 OIL TEMP

STALL WARNLANDING GEAR IND#1 STANDBY PUMP#1 OIL PRESS

LEFT BLEED AIR WARN#1 AUXILIARY TRANSFERRADOME ANTI-ICE#1 ENG AIR SCOOP HEAT

#2 DUAL FED BUS

ANN PWR#2 CHIP DETR#2 QTY IND#2 QTY WARN#2 OIL TEMPBATT CHARGE

FIRE DETRLANDING GEAR WARN#2 STANDBY PUMPX2 OIL PRESS

RIGHT BLEED AIR WARN#2 AUXILIARY TRANSFER#2 ENG AIR SCOOP HEATENG AIR SCOOP HEAT MONITOR

WSHLD WIPERSURF DEICELEFT PITOT HEATCROSSFEED#1 START CONTRPROP SYNC

#3 DUAL FED BUSA

LEFT PROP ANTI-ICELEFT FUEL VENT HEAT#1 FIREWALL VALVE#1 ICE VANE CONTR

PROP ANTI-ICE AUTOLEFT FUEL CONTR HEAT#1 PRESS WARN#1 IGNITOR CONTR

STALL WARN HEATBRAKE DEICERIGHT PITOT HEAT#2 START CONTRAUTOFEATHERHF POWER

X4 DUAL FED BUSA

RIGHT PROP ANTI-ICERIGHT FUEL VENT HEAT#2 FIREWALL VALVE#2 ICE VANE CONTRPROP GOVSCAVENGER PUMP

PROP ANTI-ICE CONTRRIGHT FUEL CONTR HEAT#2 PRESS WARN#2 IGNITOR CONTR

APO06453.2


2-60

TM 55-1510-221-10

ELEC TRIMLANDING GEAR RELAYICE LIGHTSINST INDIRECT LIGHTSTEMP CONTR

#5 DUAL FED BUS

FLAP MOTORBCN LIGHTSLANDING LIGHTSLEFT BLEED AIR CONTRPROVISIONS

PILOT TURN & SLIPSUBPANEL & CONSOLE LIGHTSRECOGNITION LIGHTSAIR COND CONTR

RUDDER BOOSTEMERG LIGHTSOVHD LIGHTSPRESS CONTRCIGAR LIGHTERAVIONICS MASTER CONTR

#6 DUAL FED BUS

FLAP CONTRFLT INST LIGHTSRIGHT BLEED AIR CONTRTAXI LIGHT

COPILOT TURN & SLIPNAV LIGHTSCABIN LIGHTSCARGO DOOR HEAT

#1 FIREWALL SHUTOFF VALVE#1 ENGINE FIRE EXTINGUISHER#1 STANDBY FUEL PUMPTRANSPONDER

HOT BATTERY BUS

CABIN LIGHTBATTERY RELAY

#2 FIREWALL SHUTOFF VALVE#2 ENGINE FIRE EXTINGUISHER#2 STANDBY FUEL PUMPCRYTO HOLD


2-61

TM 55-1510-221-10

APOO6454

Figure 2-23. Single Phase AC Electrical System

2-62

TM

55-1510-221-10

Figure 2-24. T

hree Phase A

C E

lectrical System

2-63

TM 55-1510-221-10

APOO6456

Figure 2-25. Mission Equipment DC Power System

2-64

is indicated as a percentage of total output amperagecapacity for the generating system monitored.

e. Battery Volt-Amp Meter. The mission con-trol panel (fig. 4-l), located on the right inside fuse-lage sidewall adjacent to the copilot’s seat, has a bat-tery-amperage meter that displays battery voltage onthe left side of the meter and battery current on theright side of the meter. Minimum battery voltage forengine start is 22 VDC.

f. Battery Monitor. Nickel-cadmium batteryoverheating will cause the battery charge current toincrease if thermal runaway is imminent. The air-craft has a charge-current sensor which will detect acharge current. The charge current system sensesbattery current through a shunt in the negative leadof the battery. Any time the battery charging currentexceeds approximately 7 amperes for 6 seconds orlonger, the yellow BATTERY CHARGE annuncia-tor light and the master fault caution light will illu-minate. Following a battery engine start, the cautionlight will illuminate approximately six seconds afterthe generator switch is placed in the ON position.The light will normally extinguish within two to fiveminutes, indicating that the battery is approaching afull charge. The time interval will increase if the bat-tery has a low state of charge, the battery tempera-ture is very low, or if the battery has previously beendischarged at a very low rate (i.e., battery operationof radios or lights for prolonged periods). The cau-tion light may also illuminate for short intervalsafter landing gear and/or flap operation. If the cau-tion light should illuminate during normal steady-state cruise, it indicates that conditions exist thatmay cause a battery thermal runaway. If this occurs,the battery switch shall be turned OFF and may beturned back ON only for gear and flap extension andapproach to landing. Battery may be used after a 15to 20 minute cool down period.

g. Generator Out Warning Lights. Two cau-tion/advisory annunciator panel lights inform thepilot when either generator is not delivering currentto the aircraft DC bus system. These lights are plac-arded No.1 DC GEN and No.2 DC GEN. Illumina-tion of the two MASTER CAUTION lights andeither fault light indicates that either the identifiedgenerator has failed or voltage is not sufficient tokeep it connected to the power distribution system.

The GPU shall be adjusted to28 volts maximum to preventthe aircraft.

regulate atdamage to

TM 55-1510-221-10

h. DC External Power Source. External DCpower can be applied to the aircraft through anexternal power receptacle on the underside of theright wing leading edge just outboard of the enginenacelle. The receptacle is installed inside of the wingstructure and is accessible through a hinged accesspanel. DC power is supplied through the DC exter-nal power plug and applied directly to the batterybus after passing through the external power relay.Turn off all external power while connecting thepower cable to, or removing it from, the externalpower supply receptacle. The holding coil circuit ofthe relay is energized by the external power sourcewhen the BATT switch is in the ON position. TheGPU shall be adjusted to regulate at 28 volts maxi-mum to prevent damage to the aircraft battery.

i. Security Keylock Switch, The aircraft has asecurity keylock switch (fig. 2-12) installed on theoverhead control panel, placarded OFF - ON. Theswitch is connected to the battery relay circuit andmust be ON when energizing the battery masterpower switch. The key cannot be removed from thelock when in the ON position.

j. Circuit Breakers. The overhead circuitbreaker panel (fig. 2-26) contains circuit breakers formost aircraft systems. The circuit breakers on thepanel are grouped into areas which are placarded asto the general function they protect. A DC powerdistribution panel is mounted beneath the aislewayfloor forward of the main spar. This panel containshigher current rated circuit breakers and is notaccessible to the flight crew under normal condi-tions.

2-75. AC POWER SUPPLY.

a. Single Phase AC Power Supply. AC powerfor the aircraft is supplied by inverter units, num-bered No. 1 and No.2 (fig. 2-23) which obtain opera-tional current from the DC power system. Bothinverters are rated at 750 volt-amperes and providesingle-phase output only. Each inverter provides 115volt and 26 volt, 400 Hz AC output. The invertersare protected by circuit breakers mounted on theDC power distribution panel beneath the aislewayfloor. Controls and indicators of the AC power sys-tem are located on the overhead control panel andon the caution/advisory annunciator panel.

( 1 . ) A C P o w e r WARNING/CAUTIONLights. Illumination of the two MASTER CAU-TION lights and the illumination of an annunciatorcaution light No.1 INVERTER or No.2 INVERTERindicates an inverter failure.

(2.) Instrument AC Light. A red warninglight and two MASTER WARNING lights located

2-65

TM 55-1510-221-10

Figure 2-26. Overhead Circuit Breaker Panel

on the warning annunciator panel, placarded INSTAC, will illuminate if all instrument AC bussesshould fail.

(3.) Inverter Control Switches. Twoswitches, placarded INVERTER No.1 and No.2 onthe overhead control panel (fig. 2-12) give the pilotcontrol of the single-phase AC inverters.

(4.) Volt-Frequency Meters. Two volt-frequency meters (fig. 2-12) are mounted in theoverhead control panel to provide monitoring capa-bility for both 115 VAC buses. Normal display onthe meter is shown in frequency (Hz). To read volt-age, press the button located in the lower left cornerof the meter. Normal output of the inverters will beindicated by 115 VAC and 400 Hz on the meters.

b. Three Phase AC Power Supply. Three phaseAC electrical power for operation of the inertial nav-igation system and mission avionics is supplied bytwo DC powered 3000 volt-ampere solid state threephase inverters.

(1.) Three phase inverter control switches.Two three position switches placarded #1 INV-OFF-ON-RESET and NO. 2 INV-OFF-ON-RESET,located on the mission control panel (fig. 4-1) con-trols three phase inverter operation.

(2.) Three phase volt/frequency meters.Two three phase volt/frequency meters, mounted onthe mission control panel (fig. 4-l), monitor and dis-play the voltage and frequency outputs of the threephase inverters.

(3.) Three phase loadmeters. Two threephase loadmeters, mounted on the mission controlpanel (fig. 4-l), monitors inverter output level.

(4.) Three phase AC off annunciator light.An indicator light placarded 3!zsl! AC OFF, locatedon the misson annunciator panel (fig. 4-l), indicatesa problem with one of the three phase AC powerbusses.

(5.) Three phase AC external power. Exter-nal three phase AC power for operation of the iner-tial navigation system or mission equipment, can beapplied to the aircraft through an external powerreceptacle located on the underside of the left wingleading edge just outboard of the engine nacelle (fig.2-l). The receptacle is installed inside the wingstructure and is accessible through a hinged accesspanel. The AC electrical system is automatically iso-lated from the external power source if the externalpower is over or under voltage, over or under fre-quency, or has an improper phase sequence.

2-66

TM 55-1510-221-10

(a.) External AC power annunciator (b.) External AC power controlLight. An annunciator light placarded EXT ACPWR ON, located on the mission annunciator panel

switch. A switch placarded EXT POWER-OFF-ON-

(fig. 4-1) indicates that external AC power is con-RESET, located on the mission control panel (fig.

nected to the 3 phase buses. The EXTERNAL 4-l), controls application of three phase AC power

POWER annunciator in the advisory annunciator to the aircraft.panel indicates that an AC GPU plug is mated tothe AC external power receptacle.

Section X. LIGHTING

2-76. EXTERIOR LIGHTING. located below the aisleway floor forward of the main

a. Description. Exterior lighting (fig. 2-27)consists of a navigation light on the aft end of theaft section of the vertical stabilizer, one navigationlight on the outside of each wing tip pod, two strobebeacons, one on top of the vertical stabilizer and oneon the underside of the fuselage center section, duallanding lights and a taxi light mounted on the nosegear assembly, a recognition light located in eachwing tip, and two ice lights, one light flush mountedin each nacelle, positioned to illuminate along theleading edge of each outboard wing.

b. Navigation Lights. The navigation lightsare protected by a 5-ampere circuit breaker plac-arded NAV on the overhead circuit breaker panel(fig. 2-26). Control of the lights is provided by aswitch placarded NAV-ON on the overhead controlpanel (fig. 2-l 2).

c. Strobe Beacons. The strobe beacons aredual intensity units. They are protected by a 15-ampere circuit breaker placarded BCN on the over-head circuit breaker panel (fig. 2-26). Control of thelights is provided by a switch placarded BEACON -DAY - NIGHT (fig. 2-12). Placing the switch in theDAY position will activate the high intensity whitesection of the strobe lights for greater visibility dur-ing daytime operation. Placing the switch in theNIGHT position activates the lower intensity redsection of the strobe lights.

d. Landing/Taxi Lights. Dual landing lightsand a single taxi light are mounted on the nose gearassembly. The lights are controlled by switches, plac-arded LANDING and TAXI, located in theLIGHTS section of the pilot’s subpanel. The landinglight circuit is protected by a 5-ampere circuitbreaker placarded LANDING, located on the over-head circuit breaker panel (fig. 2-26). The taxi lightcircuit is protected by a 5-ampere circuit breakerplacarded TAXI, located on the overhead circuitbreaker panel (fig. 2-26). Landing/Taxi lights areturned off when the landing gear is retracted. Thelanding lights and taxi light power circuits are pro-tected by 35-ampere and 15-ampere circuit breakers,respectively, on the DC power distribution panel

spar.

e. Ice Lights. The ice lights circuit is pro-tected by a 5-ampere circuit breaker placarded ICEon the overhead circuit breaker panel (fig. 2-26).Control of the lights is provided by a switch plac-arded ICE - ON on the overhead control panel (fig.2-12). Prolonged use during ground operation maygenerate enough heat to damage the lens.

f. Recognition Lights. A switch placardedRECOG - ON, located in the pilot’s subpanelLIGHTS section (fig. 2-6), controls the white recog-nition light in each wing tip. When requested, thissteady, bright light is used for identification in thetraffic pattern. The recognition lights circuit is pro-tected by a 7 l/2 ampere RECOG circuit breakerlocated on the overhead circuit breaker panel (fig.2-26).

2-77. INTERIOR LIGHTING.

Lighting systems are installed for use by thepilot and copilot. The lighting systems in the cockpitare provided with intensity controls on the overheadcontrol panel. A switch placarded MASTER PANELLIGHTS - ON, on the overhead control panel (fig.2-12), provides overall on-off control for all engineinstrument lights, pilot and copilot instrumentlights, overhead panel lights, console and subpanellights and the outside air temperature light.

a. Cockpit Lighting.

(1.) Flight instrument lights. Each individ-ual flight instrument contains internal lamps for illu-mination. The circuit is protected by a 7 l/2-amperecircuit breaker placarded FLT INST on the over-head circuit breaker panel (fig. 2-26). Control is pro-vided by two rheostat switches placarded PILOTINST LIGHTS - OFF - BRT and COPILOT INSTLIGHTS - OFF - BRT on the overhead controlpanel (fig. 2- 12). Turning the control clockwise fromOFF turns the lights on and increases their bril-liance.

2-67

TM 55-1510-221-10

1. Wing navigation light2. Emergency light3. Strobe beacon4. Tail navigation light5. Recognition lights6. Ice light7. Taxi light6. Landing lights

Figure 2-27. Exterior Lighting

AP 011764

2-68

TM 55-1510-221-10

(2.) Instrument indirect fights. Three lightsare mounted in the glareshield overhang along thetop edge of the instrument panel and provide overallinstrument panel illumination. The circuit is protectedby a 5-ampere circuit breaker placarded INSTINDIRECT on the overhead circuit breaker panel (fig.2-26). Control is provided by a rheostat switchplacarded INST INDIRECT LIGHTS - OFF - BRT onthe overhead control panel (fig. 2-12). Turning thecontrol clockwise from OFF turns the lights on andincreases their brilliance.

(3.) Engine instrument lights. Eachindividual engine instrument contains internal lamps forillumination. The circuit is protected by a 7 l/2- amperecircuit breaker placarded FLT INST on the overheadcircuit breaker panel (fig. 2-26). Control is provided bya rheostat switch placarded ENGINE INST LIGHTSOFF BRT on the overhead control panel (fig. 2-12).Turning the control clockwise from OFF turns the lightson and increases their brilliance.

NOTE

The floodlight is connected to the hot batterybus and will not be turned off by the batteryswitch; therefore, it must be turned OFF whenthe aircraft is shutdown to prevent dischargingthe battery.

(4.) Flood light. A single overhead floodlight is installed. It provides overall illumination of theentire cockpit area. The circuit is protected by a5-ampere circuit breaker mounted beneath the batteryand connected to the emergency battery bus. Control isprovided by a rheostat switch placarded - OVERHEADFLOODLIGHT-OFF-BRT on the overhead controlpanel (fig. 2-12). Turning the control clockwise fromOFF turns 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 1/2-amperecircuit breaker placarded OVHD on the overhead circuitbreaker panel (fig. 2-26). Control is provided by arheostat switch placarded OVERHEAD PANELLIGHTS - OFF - BRT on the overhead control panel(fig. 2-12). 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 lightedpanels, mission control panel, and pedestal extensionpanels are protected by a 7 l/2-ampere circuit breakerplacarded SUBPNL & CONSOLE on the overheadcircuit breaker panel (fig. 2-26). Control is provided bytwo rheostat switches placarded SUBPANEL orCONSOLE LIGHTS - OFF - BRT on the overhead

control panel (fig. 2- 12). Turning the control clockwisefrom OFF turns the lights on and increases theirbrilliance.

(7.) Outside air temperature light. Two postlights are mounted adjacent to the outside airtemperature gage on the left cockpit sidewall trim panel.The circuit is protected by a 71/2-ampere circuit breakerplacarded FLT INST on the overhead circuit breakerpanel (fig. 2-26). Control is provided by a pushbuttonswitch adjacent to the gage. No intensity control isprovided.

b. Cabin Lighting.

(1.) Threshold and spar cover lights. Athreshold light is installed just above floor level on theleft side of the cabin just inside the cabin door. A sparcover light is installed on the left side of the sunkenaisle immediately aft of the main spar cover. Bothcircuits are protected by a 5-ampere circuit breakermounted beneath the battery and connected to theemergency battery bus. Both lights are controlled by theswitch mounted adjacent to the threshold light. If thelights are illuminated, closing the cabin door willautomatically extinguish them.

(2.) Cabin aisle lights. Three cabin aislelights are installed in the cabin aisle. Control is providedby the CABIN LIGHTS switch on the right subpanel.Control is provided by the CABIN LIGHT switch onthe right subpanel.

(3.) Cabin spot lights. A spot light is mountedto each cabin aisle light. Each spot light is individuallycontrolled by a rheostat placarded OFF-ON-BRT on theback of the light. There is a momentary ON switch inthe center of the rheostat. Each light is capable ofproducing a red or white spotlight by turning theselector on the front of the light. To remove the lightfrom the stationary position, pull down on the light. Thelight is connected to the light housing by an 11 inchcoiled cord that extends to approximately 50 inches.

(4.) Cabin door latching mechanism light. Alight is provided to check the cabin door latchingmechanism. It is controlled by a red pushbutton switchlocated adjacent to the round observation window,which is just above the second step.

2-78. EMERGENCY LIGHTING.a. Description. An independent battery operated

lighting 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

2-69

TM 55-151 O-221 -10

placards and locate exits. An inertia switch, when ORIDE OFF - RESET - AUTO - TEST. Should thesubjected to a 2 G shock, will illuminate interior lights in system accidentally actuate, the emergency lights willthe cockpit, forward and aft cabin areas, and exterior illuminate. Placing the switch in the momentary OFFlights aft of the emergency exit and aft of the cabin door. RESET position will extinguish the lights. To test theThe battery power source is automatically recharged by system, place the switch in the TEST position. Thethe aircraft electrical system. lights should illuminate. Moving the switch to the OFF

- RESET position will turn the system off and reset it.b. Operation. An emergency lights override switch,

located on the overhead control panel (fig. 2-12), isprovided to turn the system off if it is accidentallyactuated. The switch is placarded EMERG LIGHTS

Section Xl. FLIGHT INSTRUMENTS

2-79. PITOT AND STATlC SYSTEM.

a. Description, The pitot and static system (fig.2-28) supplies static pressure to two airspeed indicators,the copilot altimeter, the air data computer (ADC), twovertical velocity indicators, and also ram air to theairspeed indicators and the ADC. This system consistsof two pitot masts (one located on each side of the lowerportion of the nose), static air pressure ports in theaircraft’s exterior skin on each side of the aft fuselage,and associated system plumbing. The pitot mast isprotected from ice formation by internal electric heatingelements.

b. Alternate Static Air Source. An alternate staticair line, 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 7for airspeed indicator and altimeter calibrationinformation when using the alternate air source.

2-80. TURN-AND-SLIP INDICATORS.

Turn-and-slip indicators are installed separately onthe pilot and copilot sides of the instrument panel (fig.2-29). These indicators are gyroscopically operated.They use DC power and are protected by 5-amperecircuit breakers placarded TURN & SLIP PILOT orCOPILOT on the overhead circuit breaker panel (fig.2-26).

2-81. AIRSPEED INDICATORS.

Airspeed indicators are installed separately on thepilot and copilot sides of the instrument panel (fig. 2-29).These indicators require no electrical power for

2-70 Change 4

operation, The indicator dials are calibrated in knotsfrom 40 to 300. A striped pointer automatically displaysthe maximum allowable airspeed at the aircraft’s present altitude.

2-82. COPILOT’S ENCODING ALTIMETER.

The copilot’s altimeter on the upper right side of theinstrument panel (fig. 2-29) is a self-contained unitconsisting of a precision pneumatic altimeter combinedwith an altitude encoder. The display face indicateswhile, simultaneously, the encoder transmits pressurealtitude information to the INS and GPS. Altitude isdisplayed by a 10,000 foot counter, a 1000 foot counter, a100 foot counter, and a single needle pointer whichindicates hundreds of feet on a circular scale in 20 footintervals. The needle pointer is also coupled to the 100foot drum counter so that both move at the same time.Below an altitude of 10,000 feet, a diagonal stripedsymbol will appear on the 10,000 foot counter. Abarometric pressure setting knob is provided to insertthe desired altimeter setting in inches Hg or millibars. IfAC power to the altitude encoder is lost, an OFF flag willappear in the upper center portion of the instrument faceto indicate that the encoder is inoperative and thesystem is not reporting altitude to ground stations.

2-83. PILOT’S ALTlMETER INDICATOR.

The pilot’s altimeter, on the upper left side of theinstrument panel (fig. 2-29), is a servoed unit undercontrol of the Air Data Computer and is part of the FlightDirector/Autopilot System. It lacks encoding capability,but displays altitude as described for the copilot’sinstrument. Operating instructions are provided inchapter 3. When the BAR0 knob is adjusted to groundsupplied instructions, the updated altitude pressure isrouted to the Air Data Computer. The ADC recomputesall data on hand, sends corrected altitude pressureinformation to the Flight Director and autopilot, sendsservo commands to correct the display on the pilot’s

TM 55-1510-221-10

Figure 2-28. Pitot and Static System

2-71

TM 55-1510-221-10

1. Airspeed indicator2. Attitude director indicator3. Flight director annunciatorpanel4. Gyro fast erect switch5. Master caution/warning annunciator6. Marker beacon dimmer control7. Marker beacon indicator lights8. Pilot’s altimeter9. Warning annunciator panel

10. Push-to-exinguish/squib OKannunciators

11. Fire pull handles12. Radar warning control panel (AN/APR-39)13. Torque indicators14. Prop Tachometers15. Turbine gas temperature indicators16. Radar signal detecting set indicator(AN/APR-39)17. Accelerometer18. RMI19. Copilot’s gyro horizon indicator20. Altitude select controller21. Copilot’s altimeter22. Vertical speed indicator

AP 011765.1

Figure 2-29. lnstrument Panel (Sheet 1 of 2)

2-72

TM 55-1510-221-10

23. PILOT SELECT annunciator24. Course indicator selector switch25. RMI select switch26. Compass #1 and #2 switch27. Microphone select switch28. Gyro INCREASE-DECREASE switch29. Gyro SLAVE-FREE switch30. Turn & slip indicator31. Compass sync annunciator32. Copilot’s horizontal situationindicator33. Cabin altitude indicator

34. Cabin rate-of-climb indicator35. HSI readout dim control36. INS control display Indicator37. TACAN range indicator38. Fuel flow gages39. Oil pressure and temp gages40. Turbine tachometers41. Radar warning control panel (AN/APR-44)42. Weather radar indicator43. Propellers synchroscope44. Propeller synchronizer switch45. Radio altimeter indicator AP 011765.246. Pilot’s horizontal situationindicator

Figure 2-29. lnstrument Panel (Sheet 2 of 2)

2-73

TM 55-1510-221-10

altimeter, and supplies altitude information to thetransponder.

2-84. VERTICAL VELOCITY INDICATORS.

Vertical velocity indicators are installed sepa-rately on the pilot and copilot sides of the instru-ment panel (fig. 2-29). They indicate the speed atwhich the aircraft ascends or descends based onchanges in atmospheric pressure. The indicator is adirect reading pressure instrument requiring no elec-trical power for operation.

2-85. ACCELEROMETER.

The accelerometer, located on the instrumentpanel registers and records positive and negative Gloads imposed on the aircraft. One hand moves inthe direction of the G load being applied while theother two, one for positive G loads and one for neg-ative G loads, follow the indicating pointer to itsmaximum travel. The recording pointers remain atthe respective maximum travel positions of the G’sbeing applied, providing a record of maximum Gloads encountered. Depressing the push-to-resetknob at the lower left comer of the instrumentallows the recording pointers to return to the normalposition.

2-86. OUTSIDE AIR TEMPERATURE (OAT)GAGE.

The outside air temperature gage, mounted out-board of the pilot’s seat, (fig. 2-8), indicates the out-side air temperature in degrees Celsius.

2-87. STANDBY MAGNETIC COMPASS.

W A R N I N G

Inaccurate indications on the standbymagnetic compass will occur while wind-shield heat and/or air conditioning isbeing used.

The standby magnetic compass is located belowthe overhead fuel management panel and to theright of the windshield divider. It may be used inthe event of failure of the compass system, or forinstrument cross check. Readings should be takenonly during level flight since errors may be intro-duced by turning or acceleration. A compass correc-tion chart indicating deviation is located on themagnetic compass.

2-88. MISCELLANEOUS INSTRUMENTS.

a. Annunciator Panels. Three annunciatorpanels are installed. One is a warning panel with redfault identification lights, and the others are caution/advisory panels with yellow and green identificationlights. The warning panel is mounted near the centerof the instrument panel below the glareshield (fig.2-29) and one caution/advisory panel is located onthe center subpanel (fig. 2-6). The mission annuncia-tor panel is located on the copilot’s sidewall. Somenormal flight operations involve indications fromthe mission control panel (fig. 4-l). Illumination ofa red warning light signifies the existence of a haz-ardous condition requiring immediate corrective

NOMENCLATURE COLOR CAUSE FOR ILLUMINATION

NO.1 FUEL PRESS RED Fuel pressure failure on left sideNO.2 FUEL PRESS RED Fuel pressure failure on right sideL BL AIR FAIL RED Left bleed air warning line has melted or failed, indicating

possible loss of No.1 engine bleed air

R BL AIR FAIL RED Right bleed air warning line has melted or failed, indicatingpossible loss of No. 2 engine bleed air

ALT WARN RED Cabin altitude exceeds 12,500 feetINST AC RED No AC power to engine instrumentsAP TRIM FAIL RED Trim inoperative or running opposite direction commanded

NO.1 CHIP DETR RED Contamination of No.1 engine oil detected

NO.2 CHIP DETR RED Contamination of No.2 engine oil detected

AP DISC RED Autopilot has disengaged.

action.

Table 2-6. Warning Annunciator Panel Legend

WARNING ANNUNCIATOR

2-74

TM 55-1510-221-10

Table 2- 7. Caution/Advisory Annunciator Panel Legend(Sheet 1 of 2)

CAUTION/ADVISORY ANNUNCIATORNOMENCLATURE COLOR CAUSE FOR ILLUMINATION

No.1 DC GEN Yellow No.1 engine generator off the line

No. 1 INVERTER Yellow No. 1 inverter inoperative

REV NOT READY Yellow Propeller levers are not in the high RPM, low pitch position,with the landing gear extended

No.2 INVERTER Yellow No.2 inverter inoperative

No.2 DC GEN Yellow No.2 engine generator off line

No.1 EXTGH DISCH Yellow No. 1 engine fire extinguisher dischargedNo.1 NAC LOW Yellow No.1 engine has 20 minutes fuel remaining at sea level, nor-

mal cruise power consumption rate

CABIN DOOR Yellow Cabin/door open or not secure

No.2 NAC LOW Yellow No.2 engine has 20 minutes fuel remaining at sea level, nor-mal cruise power consumption rate

No.2 EXTGH DISCH Yellow No.2 engine fire extinguisher discharged

No. 1 VANE FAIL Yellow No.1 engine ice vane malfunction. Ice vane has not attainedproper position

BATTERY CHARGE Yellow Excessive charge rate on battery

PROP SYNC ON Yellow Synchrophaser turned on with landing gear extended

No.2 VANE FAIL Yellow No.2 engine ice vane malfunction. Ice vane has not attainedproper position

DUCT OVERTEMP Yellow Excessive bleed air temperature in environmental heat ducts

IFF Yellow Transponder fails to reply to a valid mode 4 interogation

No.1 NO FUEL XFR Yellow Auxiliary fuel tank on side of No. 1 engine not transferring fuelinto nacelle tank

No.2 NO FUEL XFR Yellow Auxiliary fuel tank on side of No.2 engine not transferring fuelinto nacelle tank

No.1 LIP HEAT Yellow Failure of lip heat valve to conform to selected position or intransit

No.2 LIP HEAT Yellow Failure of lip heat valve to conrom to selected position or intransit

INS Yellow Inertial navigation system’s cooling fan is off or an INS mal-function that illuminates the WARN annunciator on the CDU

No.1 LIP HEAT ON Green No.1 engine air scoop heat switch is on

No.2 LIP HEAT ON Green No.2 engine air scoop heat switch is on

A/C COLD OPN Green Air conditioner is operating in cold mode, or ambient temper-atures require switching to cold mode if air conditioner opera-tion is to be continued

No.1 VANE EXT Green No. 1 ice vane extended

FUEL CROSSFEED Green Crossfeed valve open

AIR COND N, LOW Green No.2 engine RPM too low for air conditioning load

No.2 VANE EXT Green No.2 ice vane extended

No.1 IGN ON Green No.1 engine ignition/start switch on No.1 engine autoignitionswitch armed and engine torque below 20 percent

2-75

TM 55-1510-221-10

Table 2-7. Caution/Advisory Annunciator Panel Legend(Sheet 2 of 2)

CAUTION/ADVISORY ANNUNCIATORNOMENCLATURE COLOR CAUSE FOR ILLUMINATION

L BL AIR IFF Green Left environmental bleed air valve closedEXTERNAL POWER Green External power connector plugged inR BL AIR OFF Green Right environmental bleed air valve closedNo.2 IGN ON Green No.2 engine ignition/start switch on, No.2 engine autoignition

switch armed and engine torque below 20 percentNo. 1 AUTOFEATHER Green No.1 engine autofeather armed with power levers advanced

above 90% N1

No.2 AUTOFEATHER Green No.2 engine autofeather armed with power levers advancedabove 90% N1

BRAKE DEICE ON Green Brake deice system activated

a. Annunciator Panels. Three annunciatorpanels are installed. One is a warning panel with redfault identification lights, and the others are caution/advisory panels with yellow and green identificationlights. The warning panel is mounted near the centerof the instrument panel below the glareshield (fig.2-29) and one caution/advisory panel is located onthe center subpanel (fig. 2-6). The mission annuncia-tor panel is located on the copilot’s sidewall. Somenormal flight operations involve indications fromthe mission control panel (fig. 4-l). Illumination ofa red warning light signifies the existence of a haz-ardous condition requiring immediate correctiveaction. A yellow caution light signifies a conditionother than hazardous requiring pilot attention. Agreen advisory light indicates a functional situation.Table 2-6, 2-7, and 2-8 provides a list of causes forillumination of the individual annunciator lights. Infrontal view both panels present rows of small,opaque rectangular indicator lights. Word printingon each indicator identifies the monitored function,situation, or fault condition, but cannot be readuntil the light is illuminated. The bulbs of all annun-ciator panel lights are tested by activating theANNUNCIATOR TEST switch, located on the rightsubpanel near the caution/advisory panel. The sys-tem is protected by two 5-ampere circuit breakersplacarded ANN PWR and ANN IND on the over-head circuit breaker panel (fig. 2-26). The annuncia-tor system lights are dimmed when the MASTERPANEL LIGHTS switch is ON and the pilot’s flightinstrument lights are illuminated. The lights areautomatically reset to maximum brightness if:

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

(2.) T h e I N S T I N D I R E C T L I G H T Sswitch is rotated clockwise.

(3.) The MASTER PANEL LIGHTSswitch is off.

(4.) The MASTER PANEL LIGHTSswitch is ON and the PILOT INST LIGHTS switchis OFF.

(5.) Master warning light (red). A MAS-TER WARNING light is provided for both the pilotand the copilot and is located on each side of theglareshield (fig. 2-29). Any time a warning light illu-minates, the MASTER WARNING light will illumi-nate, and will stay illuminated until the MASTERWARNING light is pressed to reset the circuit. If anew condition occurs, the light will be reactivated,and the applicable annunciator panel light will illu-minate.

(6.) Master caution light (yellow). A MAS-TER CAUTION light is provided for both the pilotand copilot located adjacent to the MASTERWARNING LIGHT. Whenever a caution light illu-minates, the MASTER CAUTION will illuminate,and will stay illuminated until the condition is cor-rected and/or the MASTER CAUTION light ispressed to reset the circuit. If a new conditionoccurs, the light will be reactivated and the appro-priate annunciator panel lights will illuminate.

b. Clocks. One manually-wound 8-day clock ismounted in the center of the pilot’s control wheeland an electric digital clock is mounted in the centerof the copilot’s control wheel.

2-76

TM 55-1510-221-10

Table 2-8. Mission Control Panel Annunciator Legend

MISSION ANNUNCIATORNOMENCLATURE COLOR CAUSE FOR ILLUMINATION

MSN OVERTEMP Yellow Mission equipment is overheating.

CRYPT0 ALERT Yellow Coded messages being received.

PWR SPLY FAULT Yellow Mission power out of tolerance.

CALL Yellow Reciving transmission on VOW.

3Ø AC OFF Yellow Three phase AC power fault.

BAT FEED FAULT Yellow Ground fault detected in battery or external power line.

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.

NO INS UPDATE Yellow INS update is not in process.

TDOA OVERTEMP Yellow TDOA equipment is overheating.

LB PS OVERTEMP Yellow LB PS equipment is overheating.

TDOA FAULT Yellow TDOA system has fault.

LB PS FAULT Yellow LB PS has fault.

ELINT FAULT Yellow ELINT system has fault.

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.

MISION AC ON Green Mission AC power is on.

INS UPDATE Green INS update in process.

TDOA PWR ON Green TDOA power is on.

MISSION DC ON Green Mission DC power 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.

Section XII. SERVICING, PARKING, AND MOORING

2-89. GENERAL.

The following paragraphs include the proceduresnecessary to service the aircraft except lubrication.The lubrication requirements of the aircraft are cov-ered in the aircraft maintenance manual. Table 2-9,2-10, 2-11 and 2-12 are used for identification of

fuel, oil, etc. used to service the aircraft. The servic-ing instructions provide procedures and precautionsnecessary to service the aircraft.

2-77

TM 55-1510-221-10

DETAIL A

DETAIL C

DETAIL B

1. Air conditioning compressor2. External power receptacle3. Hand fire extinguisher4. Battery 24 VDC 5. Oxygen system filler6. Oxygen cyilnders 2 (64 cu ft bottles)

7. Electric toilet8. Fuel filler cap (typical left and right)9. Landing gear tires

10. Engine fire extinguisher11. Engine oil filler cap (typical left and right)12. Wheel brake fluid reservoir

AP011766

Figure 2-30. Servicing Locations

2-78

TM 55-1510-221-10

Table 2-9. Approved Military Fuels, Oil, Fluids, and Unit Capacities.

SYSTEM SPECIFICATION CAPACITY

Fuel MIL-T-5624 (JP4 and JP-5) 546 U.S. Gals.Engine Oil MIL-L-23699 14 U.S. Quarts per engineHydraulic Brake System MIL-H-5606 1 U.S. PintOxygen System MIL-O-27210 128 Cubic FeetToilet Chemical Monogram DG-19 3 Ounces

2-90. FUEL HANDLING PRECAUTIONS.

Table 2-2, Fuel Quantity Data, lists the quantity andcapacity of fuel tanks in the aircraft. Service the fueltanks after each flight to keep moisture out of the tanksand to keep the bladder type cells from drying out.Observe the following precautions:

WARNING

During warm weather open fuel caps slowlyto prevent being sprayed with fuel.

WARNING

When aviation gasoline is used in a turbineengine, extreme caution should be usedwhen around the combustion chamber andexhaust area to avoid cuts or abrasions. Theexhaust deposits contain lead oxide whichwill cause lead poisoning.

such as drills or buffers, in or near the aircraft duringfueling.

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

c. Carefully remove snow, water, and ice from theaircraft fuel filler cap area before removing the fuel fillercap (fig. 2-30). Remove only one aircraft filler cap at anyone time, and replace each one immediately after theservicing operation is completed.

d. Wipe all frost from fuel filler necks beforeservicing.

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

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

CAUTIONg. Observe NO SMOKING precautions.

Proper procedures for handling JP-4 andJP-5 fuel cannot be over stressed. Clean,fresh fuel shah be used and the entrance ofwater into the fuel storage or aircraft fuelsystem must be kept to a minimum.

CAUTION

When conditions permit, the aircraft shall bepositioned so that the wind will carry thefuel vapors away from all possible sources ofignition. The fuel vehicle shall be positionedto maintain a minimum distance of 10 feetfrom any part of the aircraft, whilemaintaining a minimum distance of 20 feetbetween the fueling vehicle and the fuel fillerpoint.

a. Shut off unnecessary electrical equipment on theaircraft, including radar and radar equipment. Themaster switch may be left on, to monitor fuel quantitygages, but shall not be moved during the fuelingoperation. Do not allow operation of any electrical tools,

h. Prior to transferring the fuel, insure that thehose is grounded to the aircraft.

i. Wash off spilled fuel immediately

j. Handle the fuel hose and nozzle cautiously toavoid damaging the wing skin.

k. Do not conduct fueling operations within 100 feetof energized airborne radar equipment or within 300 feetof energized ground radar equipment installations.

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

Change 4 2-79

TM 55-1510-221-10

Table 2- 10. Approved Fuels

SOURCE

US MILITARY FUELNAT0 Code No.

COMMERCIAL FUELASTM-D-l 655)American Oil Co.Atlantic RichfieldRichfield Div.B.P. TradingCaltex Petroleum Corp.Cities Service Co.Continental Oil Co.Gulf OilEXXON Co. USAMobil OilPhillips PetroleumShell OilSinclairStandard Oil Co.

PRIMARY ORSTANDARD FUEL

JP4 (MIL-T-5624)F-40 (Wide Cut Type)

JET BAmerican JP-4American JP-4Arcojet B

B.P.A.T.G.Caltex Jet-B

Conoco JP-4Gulf Jet BEXXON Turbo Fuel BMobil Jet BPhiljet JP-4Aeroshell JP-4

ALTERNATEFUEL

JP-5 (MIL-T-5624)F-44 (High Flash Type)

JET A JET A-lAmerican Type A NATO F-34American Type AArcojet A Arcojet A-1Richfield A Richfield A-1

B.P.A.T.K.Caltex Jet A-1

CITGO AConoco Jet-50 Conoco Jet-60Gulf Jet A Gulf Jet A-1EXXON A EXXON A-1Mobil Jet A Mobil Jet A-1Philjet A-50Aeroshell 640 Aeroshell 650Superjet A Superjet A-1Jet A Kerosene Jet A-1

KeroseneChevron Chevron BTexaco Texaco Avjet BUnion Oil Union JP-4Foreign Fuel NATO F-40Belgium BA-PF-2BCanada 3GP-22FDenmark JP-4 MIL-T-5624France Air 3407AGermany (West) VTL-9130-006Greece JP-4 MIL-T-5624Italy AA-M-C-1421Netherlands JP-4 MIL-T-5624Norway J P-4 M IL-T-5624Portugal JP-4 MIL-T-5624Turkey JP-4 MIL-T-5624United Kingdom (Britain) D.Eng RD 2454

Chevron A-50 Chevron A- 1Avjet B Avjet A-176 Turbine FuelNATO F-44

3-6P-24e

UTL-9130-007/UTL9130-010

AMC-143D. Eng RD 2493

D.Eng RD 2498

NOTEAnti-icing and Biocidal Additive for Commercial Turbine Engine Fuel - The fuel system icing inhibitorshall conform to MIL-L-27686. The additive provides anti-icing protection and also functions as abiocide to kill microbial growths in aircraft fuel systems. Icing inhibitor conforming to MIL-L-27686 shallbe added to commercial fuel, not containing an icing inhibitor, during refueling operations, regardlessof ambient temperatures. Refueling operations shall be accomplished in accordance with acceptedcommercial procedures.

2-80

TM 55-1510-221-10

Table 2-11. Standard Alternate and Emergency Fuels

ENGINE ARMY STANDARD FUEL ALTERNATE TYPE EMERGENCY FUELTYPE *MAX. HOURS

PT6A MIL-T-5624 MIL-T-5624 MIL-G-5572 150Grade JP-4 Grade JP-5 Any AV Gas

* Maximum operating hours with indicated fuel between engine overhauls (TBO).

2-91. FILLING FUEL TANKS. a. Army Standard Fuels. Army standard fuelis JP-4.

W A R N I N G b. Alternate Fuels, Army Alternate fuels areJP-5 and JP-8.

Prior to removing the fuel tank filler cap,the hose nozzle static ground wire shall beattached to the grounding lugs that arelocated adjacent to the filler opening.

Fill tanks as follows:

C. Emergency Fuel. Avgas is emergency fueland subject to 150 hour time limit.

2-94. USE OF FUELS.

Fuel is used as follows:

a. Attach bonding cables to aircraft.

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

Do not insert fuel nozzle completely intofuel cell due to possible damage to bottomof fuel cell. Nozzle should be supportedand inserted straight down to preventdamage to the anti-siphon valve.

c. Fill main tank before filling respective aux-iliary tanks unless less than a full fuel load isdesired.

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

e. Disconnect bonding cables from aircraft.

2-92. DRAINING MOISTURE FROM FUEL SYS-TEM.

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

2-93. FUEL TYPES.

Approved fuel types are as follows:

a. Fuel limitations. There is no special limita-tion on the use of Army standard fuel, but certainlimitations are imposed when alternate or emer-gency fuels are used. For the purpose of recording,fuel mixtures shall be identified as to the majorcomponent of the mixture, except when the mixturecontains leaded gasoline. The use of any fuels otherthan standard will be entered in the FAULTS/REMARKS column of DA Form 2408-13, AircraftMaintenance and Inspection Record, noting the typeof fuel, additives, and duration of operation.

b. Use of Kerosene Fuels. The use of kerosenefuels (JP-5 type) in turbine engines dictates the needfor observance of special precautions. Both groundstarts and air restarts at low temperature may bemore difficult due to low vapor pressure. Kerosenefuels having a freezing point of minus 40 degrees C(minus 40 degrees F) limit the maximum altitude ofa mission to 28,000 feet under standard day condi-tions.

c. Mixing of Fuels in Aircraft Tanks. Whenchanging from one type of authorized fuel toanother, for example JP-4 to JP-5, it is not necessaryto drain the aircraft fuel system before adding thenew fuel.

d. Fuel Specifications. Fuel having the sameNATO code number are interchangeable. Jet fuelsconforming to ASTM D-1655 specification may beused when MIL-T-5624 fuels are not available. Thisusually occurs during cross-country flights where air-craft using NATO F-44 (JP-5) are refueling withNATO F-40 (JP-4) or Commercial ASTM Type Bfuels. Whenever this condition occurs, the engine

2-81

TM 55-1510-221-10

operating characteristics may change in that loweroperating temperature, slower acceleration, lowerengine speed, easier starting, and shorter range maybe experienced. The reverse is true when changingfrom F-40 (JP-4) fuel to F-44 (JP-5) or CommercialASTM Type A-l fuels. Most commercial turbineengines will operate satisfactorily on either keroseneor JP-4 type fuel. The difference in specific gravitymay possibly require fuel control adjustments; if so,the recommendations of the manufacturers of theengine and airframe are to be followed.

2-95. SERVICING OIL SYSTEM

An integral oil tank occupies the cavity formedbetween the accessory gearbox housing and the com-pressor inlet case on the engine. The tank has a cali-brated oil dipstick and an oil drain plug. Avoid spill-ing 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 a satis-factory level is reached. Service oil system as fol-lows:

1. Open the access door on the upper cowlingto gain access to the oil filler cap and dip-stick.

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)for 15-20 seconds, then recheck. If over10 hours have elapsed, start the engineand run for 2 minutes, then recheck. Addoil as required. Do not overfill.

2. Remove oil filler cap.

3. Insert a clean funnel, with a screen incorpo-rated, into the filler neck.

4. Replenish with oil to within 1 quart belowMAX mark or the MAX COLD on dipstick(cold engine). Fill to MAX or MAX HOT(hot engine).

5. Check oil filler cap for damaged preformedpacking, general condition and locking.

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

2-82

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

7. Install and secure oil tiller cap.

8. Check for any oil leaks.

2-96. SERVICING HYDRAULIC BRAKE SYSTEMRESERVOIR.

1. Gain access to brake hydraulic system reser-voir.

2. Remove brake reservoir cap and till reser-voir to washer on dipstick with hydraulicfluid.

3. Install brake reservoir cap.

2-97. 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.

2-98. SERVICING THE CHEMICAL TOILET.

The toilet should be serviced during routineground maintenance of the aircraft following anyusage. The waste storage container should beremoved, emptied, its disposable plastic linerreplaced, and the container replaced in the toiletcabinet. Toilet paper, waste container plastic liners,and dry chemical deodorant packets should also beresupplied within the toilet cabinet as needed.

2-99. SERVICING THE AIR CONDITIONING SYS-TEM.

Servicing the air conditioning system consists ofchecking and maintaining the correct refrigerantlevel, compressor oil level, belt tension and condi-tion, system leak detection, and replacement of theevaporator air filters. It is imperative that the main-tenance of the air conditioning system, except for fil-ter replacement, be accomplished only by qualifiedrefrigerant system technicians. Flexible fiberglass fil-ters cover the evaporator coils and should bereplaced after 300 hours of operation. Install filtersas follows:

a. Forward Evaporator Filter Replacement:

1. Remove the access door in the nosewheel well keel under the refrigerantplumbing.

2. Pull the filter down and out of theretaining springs on the evaporatorcoil.

3. Fold the new filter to insert it throughthe access doors. The filter; must becarefully inserted between the coilassembly and the refrigerant plumbingunder the retaining springs.

4. Install the access doors.

b. Aft Evaporator Filter Replacement.

1. Remove the carpet and floor panelbehind the rear spar, and remove thecover of the evaporator plenum.

2. Remove the old filter from behind theretaining springs on the evaporatorcoil.

3. Insert the new filter between theretainer springs and the evaporatorcoil.

4. Install the plenum cover, floor panel,and carpet.

c. Anti-icing, Deicing and Defrosting Protec-tion. The aircraft is protected in subfreezing weatherby spraying the surfaces (to be covered with protec-tive covers) with defrosting fluid. Spraying defrost-ing fluid on aircraft surfaces before installing protec-tive covers will permit protective covers to beremoved with a minimum of sticking. To preventfreezing rain and snow from blowing under protec-tive covers and diluting the fluid, insure that protec-tive covers are fitted tightly. As a deicing measure,keep exposed aircraft surface wet with fluid for pro-tection against frost.

TM 55-1510-221-10

NOTE

Do not apply anti-icing, deicing anddefrosting fluid to exposed aircraft sur-faces if snow is expected. Melting snowwill dilute the defrosting fluid and form aslush mixture which will freeze in placeand become difficult to remove.

2-100. ANTI-ICING, DEICING AND DEFROSTINGTREATMENT.

Use undiluted anti-icing, deicing, and defrostingfluid (MIL-A-8243) to treat aircraft surfaces for pro-tection against freezing rain and frost. Spray aircraftsurface sufficiently to wet area, but without exces-sive drainage. A fine spray is recommended to pre-vent waste. Use diluted, hot fluid to remove iceaccumulations.

1. Remove frost or ice accumulations from air-craft surfaces by spraying with diluted anti-icing, deicing, and defrosting fluid mixed inaccordance with table 2-12.

2. Spray diluted, hot fluid in a solid stream(not over 15 gallons per minute). Thor-oughly saturate aircraft surface and removeloose ice. Keep a sufficient quantity ofdiluted, hot fluid on aircraft surface coatedwith ice to prevent liquid layer from freez-ing. Diluted, hot fluid should be sprayed ata high pressure, but not exceeding 300 PSI.

3. When facilities for heating are not availableand it is deemed necessary to remove iceaccumulations from aircraft surfaces, undi-luted defrosting fluid may be used. Sprayundiluted defrosting fluid at 15 minute

Table 2- 12. Recommended Fluid Dilution Chart

AMBIENT PERCENT PERCENT WATER FREEZING POINTTEMPERATURE DEFROSTING BY OF MIXTURE (“F)

(ºF) FLUID BY VOLUME VOLUME (APPROXIMATE)

30º and above 20 80 10º

20º 30 70 0º

10º 40 60 -15º

0º 45 55 -25º

-10º 50 50 -35º

-20º 55 45 -45º

-30º 60 40 -55º

NOTES:1. Use anti-icing and deicing fluid (MIL-A-8243) or commercial fluids.2. Heat Mixture to a temperature of 82º to 93°C (180º to 200°F).

2-83

TM 55-1510-221-10

4.

2-101.

intervals to assure complete coverage.Removal of ice accumulations using undi-luted defrosting fluid is expensive and slow.

If tires are frozen to ground, use undiluteddefrosting fluid to melt ice around tire.Move aircraft as soon as tires are free.

APPLICATION OF EXTERNAL POWER.

Before connecting the power cables fromthe external power source to the aircraft,insure that the GPU is not touching theaircraft at any point. Due to the voltagedrop in the cables, the two ground sys-tems will be of different potentials.Should they come in contact while theGPU is operating, arcing could occur.Turn off all external power while connect-ing the power cable to, or removing itfrom the external power supply recepta-cle. Be certain that the polarity of theexternal power source is the same as thatof the aircraft before it is connected. 3ini-mum GPU requirement is 400 amperescontinuous and 1800 amperes for onetenth of a second.

An external power source is often needed to sup-ply the electric current required to properly groundservice the aircraft electrical equipment and to facil-itate starting the aircraft’s engines. An external DCpower receptacle is installed on the underside of theright wing leading edge just outboard of the enginenacelle. An external AC Power receptacle is installedon the underside of the left wing leading edge justoutboard of the engine nacelle.

2-102. SERVICING OXYGEN SYSTEM.

The oxygen system furnishes breathing oxygento the pilot, copilot and first aid position Oxygencylinder location is shown in figure 2-19.

a. Oxygen System Safety Precautions.

Keep fire and heat away from oxygenequipment. Do not smoke while working

with or near oxygen equipment, and takecare not to generate sparks with carelesslyhandled tools when working on the oxy-gen system.

(1.) Keep oxygen regulators, cylinders,gages, valves, fittings, masks, and all other compo-nents of the oxygen system free of oil, grease, gaso-line, and all other readily combustible substances.The utmost care shall be exercised in servicing, han-dling, and inspecting the oxygen system.

(2.) Do not allow foreign matter to enteroxygen lines.

(3.) Never allow electrical equipment tocome in contact with the oxygen cylinder.

(4.) Never use oxygen from a cylinderwithout first reducing its pressure through a regula-tor.

b. Replenishing Oxygen System.

1. Remove oxygen access door on outsideof aircraft (fig. 2- 19).

2. Remove protective cap on oxygen sys-tem filler valve.

3. Attach oxygen hose from oxygen ser-vicing unit to filler valve.

If the oxygen system pressure is below200 PSI, do not attempt to service sys-tem. Make an entry on DA Form2408-13.

4.

5.

6.

Insure that supply cylinder shutoffvalves on the aircraft are open.

Slowly adjust the valve position so thatpressure increases at a rate not toexceed 200 PSIG per minute.

Close pressure regulating valve on oxy-gen servicing unit when pressure gageon oxygen system indicates the pres-sure obtained using the Oxygen SystemServicing Pressure Chart (fig. 2-31).

2-84

TM 55-1510-221-10

OUTSIDE AIR TEMPERATURE, °C

Figure 2-31. Oxygen System Servicing Pressure

2-85

TM 55-1510-221-10

NOTE

To compensate for loss of aircraft cylin-der pressure as the oxygen cools to ambi-ent temperature after recharging, the cyl-inder should be charged initially toapproximately 10% over prescribed pres-sure. Experience will determine what ini-tial pressure should be used to compen-sate for the subsequent pressure loss uponcooling. A small top-off will create littleheat. A complete recharge will create sub-stantial heating.

The final stabilized cylinder pressure should beadjusted for ambient temperature per figure 2-31.

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

8. Install protective cap on oxygen fillervalve.

9. Install oxygen access door.

2-103. GROUND HANDLING.

Ground handling covers all the essential infor-mation concerning movement and handling of theaircraft while on the ground. The following para-graphs give, in detail, the instructions and precau-tions necessary to accomplish ground handling func-tions. Parking, covers, ground handling and towingequipment are shown in figure 2-32.

a. General Ground Handling Procedure. Acci-dents resulting in injury to personnel and damage toequipment can be avoided or minimized by closeobservance of existing safety standard and recog-nized ground handling procedures. Carelessness orinsufficient knowledge of the aircraft or equipmentbeing handled can be fatal. The applicable technicalmanuals and pertinent directives should be studiedfor familiarization with the aircraft, its components,and the ground handling procedures applicable to it,before attempting 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 damagedor destroyed aircraft:

(1.) Keep intake air ducts free of loosearticles such as rags, tools, etc.

(2.) Stay clear of exhaust outlet areas.

(3.) During ground runup, make sure thebrakes are firmly set.

2-86

(4.) Keep area fore and aft of propellersclear of maintenance equipment.

(5.) Do not operate engines with controlsurfaces in the locked position.

(6.) Do not attempt towing or taxiing ofthe aircraft with control surfaces in the locked posi-tion.

(7.) When high winds are present, do notunlock the control surfaces until prepared to prop-erly operate them.

(8.) Do not operate engines while towingequipment is attached to the aircraft, or while theaircraft is tied down.

(9.) Check the nose wheel position. Unlessit is in the centered position, avoid operating theengines at high power settings.

(10.) Hold control surfaces in the neutralposition when the engines are being operated at highpower settings.

(11.) When moving the aircraft, do notpush on propeller deicing boots. Damage to theheating elements may result.

c. Moving Aircraft on Ground. Aircraft on theground shall be moved in accordance with the fol-lowing:

(1.) Taxiing. Taxiing shall be in accor-dance with chapter 8.

When the aircraft is being towed, a quali-fied person must be in the pilot’s seat tomaintain control by use of the brakes.When towing, do not exceed nose gearturn limits. Avoid short radius turns, andalways keep the inside or pivot wheelturning during the operation. Do not towaircraft with rudder locks installed, assevere damage to the nose steering linkagecan result. When moving the aircraftbackwards, do not apply the brakesabruptly. Tow the aircraft slowly, avoid-ing sudden stops, especially over snowy,icy, rough, soggy, or muddy terrain. Inarctic climates, the aircraft must be towedby the main gears, as an immense break-away load, resulting from ice, frozen tires,and stiffened grease in the wheel bearingsmay damage the nose gear.

Do not tow or taxi aircraft with deflatedshock struts.

(2.) Towing. Towing lugs are provided onthe upper torque knee fitting of the nose strut. Whenit is necessary to tow the aircraft with a vehicle, usethe vehicle 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/ortail by at least 15 feet. This length is required to pre-vent the aircraft from overrunning the towing vehi-cle or fouling the nose gear.

d.. Ground Handling Under Extreme WeatherConditions. Extreme weather conditions necessitateparticular care in ground handling of the aircraft. Inhot, dry, sandy, desert conditions, special attentionmust be devoted to finding a firmly packed parkingand towing area. If such areas are not available, steelmats or an equivalent solid base must be providedfor these purposes. In wet, swampy areas, care mustbe taken to avoid bogging down the aircraft. Undercold, icy, arctic conditions, additional mooring isrequired, and added precautions must be taken toavoid skidding during towing operations. The partic-ular problems to be encountered under adverseweather conditions and the special methodsdesigned to avoid damage to the aircraft are coveredby the various phases of the ground handling proce-dures included in this section of general ground han-dling instructions. (Refer to TM 55-1500-204-25/1.)

2-104. PARKING.

Parking is defined as the normal conditionunder which the aircraft will be secured while on theground. This condition may vary from the tempo-rary expedient of setting the parking brake andchocking the wheels to the more elaborate mooringprocedures described under Mooring. The propersteps for securing the aircraft must be based on thetime the aircraft will be left unattended, the aircraftweights, the expected wind direction and velocity,and the anticipated availability of ground and aircrews for mooring and/or evacuation. When practi-cal head the aircraft into the wind, especially ifstrong winds are forecast or if it will be necessary toleave the aircraft overnight. Set the parking brakeand chock the wheels securely. Following engineshutdown, position and engage the control locks.

NOTE

Cowlings and loose equipment will besuitably secured at all times when left inan unattended condition.

TM 55-1510-221-10

a. The parking brake system for the aircraftincorporates two lever-type valves, one for eachwheel brake. Both valves are closed simultaneouslyby pulling out the parking brake handle. Operate theparking brake as follows:

1. Depress both brakes.

2. Pull parking brake handle out. Thiswill cause the parking brake valves tolock the hydraulic fluid under pressurein the parking brake system, therebyretaining braking action.

3. Release brake pedals.

Do not set parking brakes when thebrakes are hot during freezing ambienttemperatures. Allow brakes to cool beforesetting parking brakes.

4. To release the parking brakes push inon the parking brake handle.

b. The control lock (fig. 2-18) holds theengine and propeller control levers in a secure posi-tion. It also holds the elevators and rudder at neutralposition and the ailerons in a staggered attitude, oneslightly "up" and the other slightly "down". Installthe control locks as follows:

1.

2.

3.

4.

With engine and propeller controllevers in secure position, slide lockonto control pedestal to prevent opera-tion of levers.

Install elevator and aileron lockpin ver-tically through pilot’s control columnto lock control wheel.

Install rudder lock pin through flapperdoor forward of pilot’s seat, makingsure rudder is in neutral position.

Reverse steps 1 through 3 above toremove control lock. Store control lock.

2-105. INSTALLATION OF PROTECTIVE COV-ERS.

The crew will insure that the aircraft protectivecovers are installed.

2-106. MOORING.

The aircraft is moored to insure its immovabil-ity, protection, and security under various weather

2-87

TM 55-1510-221-10

Figure 2-32. Parking, Covers, Ground Handling, and Towing Equipment

2-88

TM 55-1510-221-10

conditions. The following paragraphs give, in detail,the instructions for proper mooring of the aircraft.

a. Mooring Provisions. Mooring points (fig.2-33) are provided beneath the wings and tail. Addi-tional mooring cables may be attached to each land-ing gear. General mooring equipment and proce-dures necessary to moor the aircraft, in addition tothe following, are given in TM 55-1500-204-25/1.

(1.) Use mooring cables of 1/4 inch diam-eter aircraft cable and clamp (clip-wire rope), chainor rope 3/8 inch diameter or larger. Length of thecable or rope will be dependent upon existing cir-cumstances. Allow sufficient slack in ropes, chains,or cable to compensate for tightening action due tomoisture absorption of rope or thermal contractionof cable or chain. Do not use slip knots. Use bowlineknots to secure aircraft to mooring stakes.

(2.) Chock the wheels.

b. Mooring Procedures for High Winds. Struc-tural damage can occur from high velocity winds;therefore, if at all possible, the aircraft should bemoved to a safe weather area when winds above 75knots are expected. Moored aircraft condition isshown in figure 2-33. If aircraft must be secured usethe following steps:

1. After aircraft is properly located, placenose wheel in centered position. Headaircraft into the wind, or as nearly soas is possible within limits determinedby locations of fixed mooring rings.When necessary, a 45 degree variationof direction is considered to be satis-factory. Locate each aircraft at slightlymore than wing span distance from allother aircraft. Position nose mooringpoint approximately 3 to 5 feet down-wind from ground mooring anchors.

2. Deflate nose wheel shock strut towithin 3/4 inch of its fully deflatedposition.

3. Fill all fuel tanks to capacity, if timepermits.

4. Place wheel chocks fore and aft ofmain gear wheels and nose wheel. Tieeach pair of chocks together with ropeor join together with wooden cleatsnailed to chocks on either side ofwheels. Tie ice grip chocks togetherwith rope. Use sandbags in lieu ofchocks when aircraft is moored on steelmats. Set parking brake as applicable.

5. Accomplish aircraft tiedown by utiliz-ing mooring points shown in figure

2-33. Make tiedown with 1/4 inch air-craft cable, using two wire rope clips orbolts, and a chain tested for a 3000pound pull. Attach tiedowns so as toremove all slack. (Use a 3/4-inch orlarger manila rope if cable or chain tie-down is not available.) If rope is usedfor tiedown, use anti-slip knots, such asbowline knot, rather than slip knots. inthe event tiedown rings are not avail-able on hard surfaced areas, move air-craft to an area where portable tie-downs can be used. Locate anchor rodsat point shown in figure 2-33. Whenanchor kits are not available, use metalstakes or deadman type anchors, pro-viding they can successfully sustain aminimum pull of 3000 pounds.

6. In event nose position tiedown is con-sidered to be of doubtful security dueto existing soil condition, drive addi-tional anchor rods at nose tiedownposition. Place padded work stand orother suitable support under the aftfuselage tiedown position and secure.

7. Place control surfaces in locked posi-tion and trim tab controls in neutralposition. Place wing flaps in up posi-tion.

8. The requirements for dust excluders,protective covers, and taping of open-ings will be left to the discretion of theresponsible maintenance officer or thepilot of the transient aircraft (fig. 2-32).

9. Secure propellers to prevent windmill-ing (fig. 2-32).

10. Disconnect battery.

11. During typhoon or hurricane wind con-ditions, mooring security can be fur-ther increased by placing sandbagsalong the wings to break up the aerody-namic flow of air over the wing,thereby reducing the lift being appliedagainst the mooring by the wind. Thestorm appears to pass two times, eachtime with a different wind direction.This will necessitate turning the air-craft after the first passing. sli.Afterhigh winds, inspect aircraft for visiblesigns of structural damage and for evi-dence of damage from flying objects.Service nose shock strut and reconnectbattery.

2-89

TM 55-1510-221-10

NOTE

IF STRONG WINDS ARE ANTICIPATED OR AIRCRAFT ISTO BE LEFT UNATTENDED, PROPELLER RESTRAINT,PITOT MAST, AND INTAKE COVERS MUST BE IN-STALLED, AND THE FLIGHT CONTROLS LOCKENGAGED.

BEFORE TOWING, THE PROPELLER RESTRAINT MUSTBE INSTALLED WITH ONE PROPELLER BLADE IN THEDOWN POSITION AS SHOWN.

THE USE OF DOUBLE OR SINGLE MOORING POINTSFOR NOSE AND/OR WING TIEDOWNS IS DETERMINEDBY LOCAL OPTION DEPENDING ON TYPE AND AVAILA-BILITY OF AIRCRAFT SECURING EQUIPMENT.

USE ROPE ONLY (NYLON TYPE IF AVAILABLE) FORNOSE TIEDOWN (DETAIL A). ATTACH ROPE(S) TO AIR-CRAFT AND GROUND MOORING POINTS IN A MANNERTHAT WILL PREVENT ROPE DAMAGE TO AIRCRAFTCOMPONENTS.

Figure 2-33. Mooring the Aircraft

2-90

TM 55-1510-221-10

CHAPTER 3

AVIONICS

Section I. GENERAL

3-1. INTRODUCTION.

Except for mission avionics, this chapter coversall avionics equipment installed in the RC-12H air-craft. It provides a brief description of equipmentcovered, the technical characteristics and locations.It covers systems and controls and provides theproper techniques and procedures to be employedwhen operating the equipment. For more detailedoperational information consult the vendor manualsthat accompany the aircraft loose tools.

3-2. AVIONICS EQUIPMENT CONFIGURATION.

The aircraft avionics covered consists of threegroups of electronic equipment. The communicationgroup consists of the interphone, UHF command,backup VOW, VHF/AM-FM, VHF command andHF command systems. The navigation group pro-vides the pilot and copilot with the instrumentationrequired to establish and maintain an accurate flightcourse and position, and to make an approach oninstruments under Instrument Meteorological Con-ditions (IMC). The navigation group includes equip-ment for determining altitude, attitude, position,destination, range and bearing, heading reference,groundspeed, and drift angle. The transponder andradar group includes an identification, position,emergency tracking system, a radar system to locatepotentially dangerous weather areas, and a radar sys-tem to differentiate between friendly and unfriendlysearch radar.

battery, left and right generators, and externalpower. Power is routed through two 50-ampere cir-cuit breakers to the avionics power relay which iscontrolled by the AVIONICS MASTER POWERswitch on the overhead control panel (fig. 2-12).Individual system circuit breakers and the associ-ated avionics busses are shown in fig. 2-22. With theswitch in the ON (forward) position, the avionicspower relay is de-energized and power is appliedthrough both the AVIONICS MASTER POWERNo. 1 and No.2 circuit breakers to the individual avi-onics circuit breakers on the overhead circuitbreaker panel (fig. 2-26). In the OFF (aft) position,the relay is energized and power is removed fromavionics equipment. When external power is appliedto the aircraft, the avionics power relay is normallyenergized, removing power from the avionics equip-ment. To apply external power to the avionicsequipment, move the AVIONICS MASTERPOWER switch to the EXT PWR position. This willde-energize the avionics power relay and allowpower to be applied to the avionics equipment.

NOTE

b. Single-Phase AC Power. AC power for theavionics equipment is provided by two inverters.The inverters supply 115-volt and 26-volt single-phase AC power when operated by the INVERTERNo. 1 or No.2 switches (fig. 2-12). Either inverter iscapable of powering all avionics equipment requir-ing AC power. AC power from the inverters isrouted through fuses in the nose avionics compart-ment.

All avionics equipment require a 3-min-ute warmup period. The weather radarhas an automatic time delay of 3 to 4minutes.

3-3. POWER SOURCE.

a. DC Power. DC power for the avionicsequipment is provided by four sources: the aircraft

c. Three-Phase AC Power. Three phase ACelectrical power for operation of the inertial naviga-tion system and mission avionics is supplied by twoDC powered 3000 volt-ampere solid state threephase inverters. The three phase inverters are con-trolled by two three-position switches located on themission control panel (fig. 4-1) placarded No. 1 INV- OFF - ON - RESET and #2 INV - OFF - ON -RESET.

3-1

TM 55-1510-221-10

3-4. DESCRIPTION.

The communications equipment group consistsof an interphone system connected to individualaudio control panels for the pilot and copilot whichinterface with VHF, UHF, BU VOW, VHF AM-FMand HF communication units.

3-5. MICROPHONES, SWITCHES AND JACKS.

Boom and oxygen mask microphones can be uti-lized in the aircraft.

a. Microphone Switches. The pilot and copilotare provided with individual microphone controlswitches, placarded INTPH-XMIT-MIC, attached torespective control wheels. A foot-actuated micro-phone switch is also positioned on the floorboardsforward of each pilot’s seat.

b. Controls and Functions.

(1.) Microphone control wheel switches(fig. 2-17). Keys selected facility.

(a . ) INTPH (depressed to f i rs tdetent). Keys interphone facility, disregards positionof transmitter selector switch.

(b.) XMIT (depressed full down).Keys facility selected by transmitter select switch.

(2.) Floorboard microphone switches. Con-trols connection of selected microphone to audiosystem.

(a.) Held depressed. Connectsselected microphone to audio system.

(b.) Released. Disconnects selectedmicrophone from audio system.

c. Microphone jack selector switches. T w oswitches, placarded MIC HEADSET - OXYGENMASK, are located on the extreme left and extremeright of the instrument panel (fig. 2-29). Theseswitches provide a means of selecting whether theheadset microphone jack or the oxygen mask micro-phone jack is connected to the audio system.

d. Controls and Functions.

(I.) MIC HEADSET - OXYGEN MASKswitch. Selects microphone jack to connect to audiosystem.

(a.) MIC HEADSET. Connects head-set microphone to audio system.

3-2

Section II. COMMUNICATIONS

(b.) OXYGEN MASK. Connectsmicrophone in oxygen mask to audio system.

3-6. AUDIO CONTROL PANELS.

a. Description. Separate but identical audiocontrol panels (fig. 3-1), serve the pilot and copilot.The controls and switches of each panel provide theuser with a means of selecting desired reception andtransmission sources, and also a means to controlthe volume of audio signals received for interphone,communication and navigation systems. The userselects between the UHF, VHF, BU VOW, VHFAM-FM and HF transceivers. The audio controlpanels are protected by respective 2-ampere AUDIOPILOT and AUDIO COPILOT circuit breakerslocated on the overhead circuit breaker panel (fig.2-26).


(I.) Master VOL control. Controls side-tone volume to headset. Also serves as final volumeadjustment for received audio from any sourcebefore admission to headset.

(2.) Radios audio monitor controls. Eachis combination rotary control and on-off push-pullswitch, permitting both receiver selection and vol-ume adjustment.

(a.) No. 1. On connects user’s headsetto audio from VHF-AM transceiver No. 1.

(b.) No.2. On connects user’s headsetto audio from the VHF/AM/FM transceiver.

(c.) No.3. On connects user’s headsetto audio from No. 1 UHF transceiver in use.

(d.) No.4. On connects user’s headsetto audio from HF or VOW transceivers.

(e.) No.5 On connects user’s headsetto audio from No. 2 UHF (BU VOW) transceiver.

(3.) NAV receiver audio monitor controls.Combination volume control and “ON-OFF“switches for NAV receivers.

(a.) NAV-A. On connects user’s head-set to audio from VOR-1, VOR-2 or marker beaconset in use.

(b.) NAV-B. On connects user’s head-set to audio from TACAN or ADF set in use.

(4.) Microphone impedence select switch.Two-position, thumb-actuated switch. Enables selec-

TM 55-1510-221-10

Figure 3-1. Audio Control Panel (Typical Pilot, Copilot)

tion of interface circuit with best impedence matchto microphone used.

(a.) The impedance of MIC 1 posi-tion is 5 Ohms.

(b.) The impedance of MIC 2 posi-tion is 150 Ohms.

(5.) Transmitter-interphone selector switch.Connects microphone and headset to selected radiotransmitter or interphone line routing receivedaudio to headset. Bypasses control of respectivereceiver audio switch.

(a.) PVT. Position not used.

intercom.(b.) ICS. Activates pilot-to-copilot

(c.) No.1. Permits audio receptionfrom VHF-AM No. 1 transceiver. Routes key andmicrophone signals to VHF-AM No. 1 transceiver.

(d.) No.2. Permits audio receptionfrom VHF/AM/FM transceiver. Routes key andmicrophone signals to VHF/AM/FM transceiver.

(e.) No.3. Permits audio receptionfrom No. 1 UHF transceiver. Routes key and micsignals to No. 1 UHF transceiver.

(f.) No.4. Permits audio receptionfrom HF or VOW transceivers. Routes key or micro-phone signals to transceiver.

(g.) No.5 Permits audio receptionfrom No. 2 UHF (BU VOW). Routes key and andmicrophone signals to transceiver.

(6.) ICS select switch. Controls activationof microphones.

(a.) HOT MIC. Admits speech tointerphone system without need to key selectedmicrophone.

(b.) NORM. Blocks speech frominterphone system unless selected microphone iskeyed.

(c.) ICS OFF. Deactivates inter-phone system.

c. Normal Operation.

(1.) Turn-on procedure: Both audio con-trol panels are activated when electrical power isapplied to aircraft.

3-3

TM 55-1510-221-10

NOTE

It is presumed the AVIONICS MASTERPOWER switch is ON, and that normallyused avionics circuit breakers remain set.The circuit breakers of routinely used avi-onic systems are normally left set.

(2.) Receiver operating procedure:

1. Receiver audio switches (audiocontrol panel) - As required.

2. Master volume control (audiocontrol panel) - Do not use. (Ad-just volume control of systembeing used.)

NOTE

Audio select switches and volume con-trols are routinely left in positions of nor-mal use.

3. Move each receiver audio switchON then OFF, separately, to ver-ify audio presence in headphonesfor each system. Adjust volume.

(3.) Transmitter operating procedure:

1. Transmitter-interphone selectorswitch (audio control panel) - Setfor transceiver desired.

2. Microphone jack selector switch(instrument panel, fig. 2-29) - Asdesired.

3. Control wheel microphone switch(control wheel) - XMIT.

4. Microphone switch (floorboard) -Depress to transmit.

(4.) Intercommunication procedure:

1. Transmitter-interphone selectorswitch (audio control panel) - ICS.

2. Microphone jack selector switch(instrument panel, fig. 2-29) - Asdesired.

3. ICS select switch (audio controlpanel, fig. 3-1) - As desired(NORM-HOT MIC-ICS OFF).

4. If HOT MIC is selected - Talkwhen ready.

5. If NORM microphone is selected- Depress microphone switch andtransmit.

6. If ICS OFF is selected intercomfunction is switched off.

7. Volume control (selected trans-ceiver) - Set for comfort.

d. Emergency Operation. Not applicable.

e. Shutdown Procedure.

1. AVIONICS MASTER POWERswitch (overhead control panel,fig. 2-12) - OFF.

2. Leave controls and circuit break-ers positioned for normal opera-tion.

3-7. MARKER BEACON AUDIO CONTROLPANEL (FIG. 3-2).

a. Description. The marker beacon audio con-trol panel, located on the pedestal extension (fig.2-7) allows the pilot or copilot to control the volumeof the marker beacon (MKR BCN). It also has con-trols for the selection of ADF voice or range filtersand MKR BCN HI-LO sensitivity.


(1.) ADF Filter Switches. Controls selectedADF filter.

(a.) FILTER V-OFF switch. Selectsfilter to block voice transmissions from ADF groundstation.

(b.) FILTER R-OFF switch. Selectsfilter to block range transmissions from ADF groundstation.

(2.) MKR BCN volume control. Adjustsvolume of marker beacon radio signals received.

(3.) MKR BCN HI-LO switch. Selects sen-sitivity of marker beacon receiver.

3-8. UHF COMMAND SET (AN/ARC-184).

a. Description. The UHF command set is aline-of-sight radio transceiver which provides trans-mission and reception of amplitude modulated(AM) signals in the ultra high frequency range of225.000 to 399.975 MHz for a distance range ofapproximately 50 miles. Channel selection is spacedat 0.025 MHz. A separate receiver is incorporated toprovide monitoring capability for the UHF guardfrequency (243.0 MHz). UHF audio output isapplied to the audio panel where it is routed to theheadsets.

3-4

TM 55-1510-221-10

Figure 3-2. Marker Beacon Audio Control Panel

NOTE

The PRESET channel selector and man-ual frequency selectors are inoperativewhen the mode selector is set to GUARDposition. The receiver-transmitter will beset to the emergency frequency only.

The transmitter and receiver sections of theUHF unit operate independently but share the samepower supply and frequency control circuits. Sepa-rate cables route transmit and receive signals to theirrespective receiver/transmitter.

Complete provisions only are installed for aTSEC/KY-28 voice security device to be located onthe LH forward avionics rack behind the pilot. TheUHF command set is protected by the 71/2 ampereUHF circuit breaker on the overhead circuit breakerpanel (fig. 2-26). Figure 3-3 illustrates the UHFcommand set. The associated blade type antenna isshown in figure 2-1.

b. Controls and Functions. UHF control panel(fig. 3-3):

(1.) Manual frequency selector/indicator(hundreds). Selects and indicates hundreds digit offrequency (2 or 3) in MHz.

(2.) Manual frequency selector/indicator(tens). Selects and indicates tens digit of frequency(0 through 9) in MHz.

(3.) Manual frequency selector/indicator(units). Selects and indicates units digit of frequency(0 through 9) in MHz.

(4.) Preset channel indicator. Displays pre-set channel.

(5.) Manual frequency selector/indicator(tenths). Selects and indicates tenths digit of fre-quency (0 through 9) in MHz.

(6.) Preset channel selector. Selects one of20 preset channel frequencies.

(7.) Manual frequency selector (hundredthsand thousandths). Selects hundredths and thou-sandths digits of frequency (00, 25, 50, or 75) inMHz.

(a.) Mode selector. Selects operatingmode and method of frequency selection.

1. MANUAL. Enables themanual selection of any one of 7,000 frequencies.

2. PRESET. Enables selectionof any one of 20 preset channels.

3-5

TM 55-1510-221-10

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

3 . GUARD. Selection auto-matically tunes the main receiver and transmitter tothe guard frequency and the guard receiver isenabled.

(d.) ADF. Not used.


(b.) SOUELCH switch. Turns main(1.) Turn on procedure:

receiver squelch on or off.

(c.) VOL control. adjusts volume.NOTE

(8.) TONE pushbutton. When pressed,transmits a 1,020 Hz tone on the selected frequency.

It is presumed aircraft power is on andnormally used avionic circuit breakersremain depressed.

(9.) Function selector. Selects operatingfunction.

(a.) OFF. Turns set off.

(b.) MAIN. Selects normal transmis-sion with reception on main receiver.

(c.) BOTH. Selects normal transmis-sion with reception on both the main receiver andthe guard frequency receiver.

1. Avionics master power switch(overhead panel, fig. 2-12) - ON.

2. Function select switch (UHF con-trol panel, fig. 3-3) -MAIN orBOTH position, as required.

3-6

NOTE

If function selector is at MAIN setting,only the normal UHF communicationswill be received. If selector is at BOTHposition, emergency communications onthe guard channel and normal UHF com-munications will both be received.


1. UHF audio monitor switch (No.3,audio control panel) -ON, ortransmitter-interphone selectorswitch (audio control panel) -No.3 position.

2. Volume control (UHF controlpanel) - Mid position.

(3.) To use preset frequency (UHF controlpanel):

1.

2.

(4 . ) Tocontrol panel):

1.

2.

Mode selector switch - PRESETposition.

Preset channel selector switch -Rotate to desired channel.

use non-preset frequency (UHF

Mode selector switch - MANUALposition.

Manual frequency selectors (5) -Rotate each knob to set desiredfrequency digits.

NOTE

The PRESET channel selector and man-ual frequency selectors are inoperativewhen the mode selector switch is set tothe GUARD position.

3. Volume - Adjust.

NOTE

To adjust volume when audio is not beingreceived, turn squelch switch OFF, adjustvolume for comfortable noise level, thenturn squelch switch ON.

4. Squelch - As desired.


1. Transmitter-interphone selector(audio panel control panel, fig.3-1) - No.3 position.

TM 55-1510-221-10

2. UHF control panel (fig. 3-3) - Setrequired frequency using eitherPRESET CHAN control or MAN-UAL frequency select controls.

3. Microphone jack selector switch(instrument panel, fig.2-29) - Asdesired.

4. Microphone switch - Depress totransmit.

(6.) Shutdown procedure: Function selec-tor switch (UHF control panel, fig. 3-3) -OFF.

d. UHF Command Set Voice Security Opera-tion (KY-28).

NOTE

Disregard operating procedures involvingthe voice security control-indicator if unitis not installed.

(I.) Turn-on procedure:

1. Power switch (Voice Securitypanel, fig 3-6) - ON.

2. Function switch (UHF controlpanel) - BOTH.


1. Mode selector switch (UHF con-trol panel, fig. 3-3) - As required.

2. Transmitter-interphone selectorswitch (audio control panel, fig.3-1) - No.3 position, or No.3radio monitor control - On.

3. Set required frequency using pre-set channel control or manual fre-quency selector.

4. Volume control - Adjust.

NOTE

To adjust volume when radio signals arenot being received, turn squelch switchOFF, adjust volume for comfortable noiselevel, then turn squelch disable switchON.

5. Squelch switch - As required.

(3.) Transmitter operating procedure(PLAIN):

1. Transmitter-interphone selectorswitch (audio control panel) -No.3 position.

3-7

TM 55-1510-221-10

2. Plain/cipher switch (voice securitycontrol panel) -PLAIN.

3. Microphone switch - Press.

(4.) Transmitter operating procedure (CI-PHER):

1. Transmitter-interphone selectorswitch (audio control panel) -No.3 position.

2. Plain/cipher switch (voice securitycontrol panel, fig. 3-6) - CIPHER.(CIPHER indicator will be illumi-nated as long as switch is inCIPHER position.)

3. RE-X/REG switch (voice securitycontrol panel) - REG.

4. Microphone switch - Pressmomentarily (interrupted tonefrom voice security unit shouldno longer be heard).

NOTE

No traffic will be passed if the interruptedtone is still heard after pressing andreleasing the microphone switch.

5. Microphone switch - Press (do nottalk). Wait until beep is heardthen speak into microphone.

(5.) Shutdown procedure:

1. Function selector switch (UHFcontrol panel)- OFF.

2. Power switch (Voice security con-trol panel) - OFF.

tion:e. UHF Command Set - Emergency Opera-

NOTE

Transmission on emergency frequencies(guard channels) is restricted to emergen-cies only. An emergency frequency of 121.500 MHz is also available on the VHFcommand radio set.

1. Transmitter-interphone selector switch(audio control panel) - No.3 position.

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


3-8A. UHF COMMAND SET (AN/ARC-164).

a. Description. The UHF command set is a line-of-sight radio transceiver which provides transmis-sion and reception of amplitude modulated (AM) sig-nals in the ultra high frequency range of 225.000 to399.975 MHz for a distance range of approximately50 miles. Channel selection is spaced at 0.025 MHz.A separate receiver is incorporated to provide moni-toring capability for the UHF guard frequency (243.0MHz). UHF audio output is applied to the audiopanel where it is routed to the headsets.

NOTEThe PRESET channel selector and manualfrequency selectors are inoperative whenthe mode selector is set to GUARD posi-tion. The receiver-transmitter will be setto the emergency frequency only.

Existing capabilities of the HAVE QUICK modi-tied radio are preserved to the maximum extent pos-sible when it is operated in the normal (non-hopping)mode. No new procedures are required for normalradio operation.

To operate in the AJ mode, the radio must first beinitialized. This initialization requires the setting oftwo control entries into the radio, Word-of-Day(WOD) and Time-of-Day (TOD). The WOD definesthe choice of frequency hopping pattern for the day.The WOD choice is a managerial function and thesame WOD may be used for one or more days. TheTOD must be loaded into the clock contained withinthe radio.

The transmitter and receiver sections of the UHFunit operate independently, but share the samepower supply and frequency control circuits. Sepa-rate cables, route transmit and receive signals to theirrespective receiver/transmitter.

The UHF command set is protected by the 7 1/2-ampere UHF circuit breaker in the overhead circuitbreaker panel (fig. 2-6). Figure 3-2 illustrates theUHF command set. The associated blade typeantenna is shown in figure 2-1.

b. Controls and Functions. UHF control panel(fig. 3-2):

(1) Manual frequency selector/indicator(hundreds). Selects and indicates hundreds digit offrequency (2 or 3) in MHz.

3-8 Change 4

TM 55-1510-221-10

(2) Manual frequency selector/indicator(tens). Selects and indicates tens digit of frequency (0through 9) in MHz.

(3) Manual frequency selector/indicator(units). Selects and indicates units digit of frequency(0 through 9) in MHz.

(4) Preset channel indicator. Displays presetchannel.

(5) Manual frequency selector/indicator(tenths). Selects and indicates tenths digit of fre-quency (0 through 9) in MHz.

(6) Preset channel selector. Selects one of 20preset channel frequencies.

(7) Manual frequency selector (hundredthsand thousandths). Selects hundredths and thou-sandths digits of frequency (00, 25, 50, or 75) inMHz.

(8) Mode selector. Selects operating modeand method of frequency selection.

(a) MANUAL. Enables the manual selec-tion of any one of 7,000 frequencies.

(b) PRESET. Enables selection of any oneof 20 preset channels.

(c) GUARD. Selection automaticallytunes the main receiver and transmitter to the guardfrequency and the guard receiver is enabled.

(9) SQUELCH switch. Turns main receiversquelch on or off.

(10) VOL control. Adjusts volume.

(11) TONE pushbutton. When pressed, trans-nits a 1,020 Hz tone on the selected frequency.

(12) Function selector. Selects operating‘unction.

(a) OFF. Turns set off.

(b) MAIN. Selects normal transmissionwith reception on main receiver.

(c) BOTH. Selects normal transmissionwith reception on both the main receiver and theguard frequency receiver.

(d) ADF. Not used.


(1) Turn on procedure:

NOTEIt is presumed aircraft power is on andnormally used avionic circuit breakersremain depressed.

1. Avionics master power switch - ON.

2. Function select switch - MAIN orBOTH position, as required.

NOTEIf function selector is at MAIN setting,only the normal UHF communicationswill be received. If selector is at BOTHposition, emergency communications onthe guard channel and normal UHF com-munications will both be received.

(2) Receiver operating procedure:

1. Transmitter-interphone selectorswitch - No. 3 position.

2. UHF audio monitor switch - ON, No.3 position.

3. Volume control - Mid position.

(3) To use preset frequency:

1. Mode selector switch - PRESET posi-tion.

2. Preset channel selector switch -Rotate to desired channel.

(4) To use non-preset frequency:

1. Mode selector switch - MANUALposition.

2. Manual frequency selectors (5) -Rotate each knob to set desired fre-quency digits.

NOTEThe PRESET channel selector and manualfrequency selectors are inoperative whenthe mode selector switch is set to theGUARD position.

3. Volume - Adjust.

Change 4 3-8.1

TM 55-1510-221-10

Figure 3-2. UHF Control Panel (AN/ARC-164)

NOTETo adjust volume when audio is not beingreceived, turn squelch switch OFF, adjustvolume for comfortable noise level, thenturn squelch switch ON.

4. Squelch - As desired.

(5) Transmitter operating procedure:

1. Transmitter-interphone selector - No.3 position.

2. UHF control panel - Set required fre-quency using either PRESET CHANcontrol or MANUAL frequency selectcontrols.

3. Microphone jack selector switch - Asdesired.

4. Microphone switch - Depress to trans-mit.

(6) Shutdown procedure: Function selectorswitch (fig. 3-2) - OFF.

3-9. VOICE ORDER WIRE (AN/ARC-194).

A radio set identical in type and performance tothe UHF command set (fig. 3-3) is located in thepedestal, to serve as voice order wire. This set pro-vides the pilot and copilot with secure 2-way voicecommunications. Complete provisions only are pro-vided for a KY-58 voice security device. The voiceorder wire set is protected by a 7 1/2 ampere VOWcircuit breaker on the overhead circuit breaker panel(fig. 2-26). The voice order wire shares an antennamounted on the aircraft belly with the transponder(lower antenna, fig 2-1).

3-8.2 Change 4

TM 55-1510-221-10

3-10. VHF-AM COMMUNlCATlONS (VHF-206).

a. Description. VHF-AM communicationsprovide transmission and reception of amplitudemodulated signals in the very high frequency rangeof 116.000 to 151.975 MHz for a range of approxi-mately 50 miles, varying with altitude. A dual headcontrol panel (fig. 3-4) is mounted on the pedestalextension, accessible to both the pilot and copilot.The panel. provides two sets of control indicators,frequency indicators, frequency select knobs, a sin-gle volume control, and a single selector switch forquick frequency changing. Transmission audio isrouted by pilot and copilot No.1 transmitter selectorswitches located on the audio control panel (fig.3-1). Received audio is routed by pilot and copilotNo. 1 receiver audio switches (fig. 3-1), to the respec-tive headsets. The VHF radio is protected by a 10-ampere VHF circuit breaker on the overhead circuitbreaker panel (fig. 2-26). The associated antenna isshown in figure 2-1.

b. Controls/Indicators and Functions.

(1.) Frequency indicator. Indicates setoperating frequency.

(2.) Control frequency indicators. Indicatesfrequency selected (left or right active).

(3.) COMM TEST switch. Overrides auto-matic squelch circuit.

(4.) Frequency selectors. Select desired setoperating frequency.

(5.) TRANS switch. Selects right or leftfrequency control selectors.

(6.) VOL-OFF control. Adjusts volume ofreceived audio, turns set ON or OFF.

c. VHF-AM Set - Normal Operation.

(1.) Turn-on procedure: Volume control -Turn clockwise (ON).

Change 4 3-8.3/(3-8.4 blank)

TM 55-1510-221-10

Figure 3-4. VHF-AM Control Panel (VHF-20B)


1. Transmitter-interphone selectorswitch (audio control panel, fig.3-1) - No.1 position, or radiomonitor control No.1 -ON.

2. Frequency selector - Set desiredfrequency.

3. Volume control - As required.


1. Transmitter-inter-phone selectorswitch (audio control panel, fig.3-1) - No.1 position.


(4.) Shutdown procedure: Volume control- Turn counterclockwise (OFF).

d. VHF-AM Set Emergency Operation.

NOTE

Transmission on emergency frequency(121.500 MHz) is restricted to emergen-cies only. Emergency frequency 243.000MHz (guard channel) is also available onthe UHF command radio.

1. Transmitter-interphone selectorswitch (audio control panel, fig.3-1) - No. 1 position.

2. Frequency selector (VHF controlpanel, fig. 3-4) -121.500 MHz(emergency frequency).


3-11. VHF AM-FM COMMAND SET (AN/ARC-199).

a. Description. The VHF AM-FM CommandSet provides for normal and secure 2-way AM voicecommunication in the very high frequency range of116.000 to 151.975 MHz and FM voice communica-tion in the 30.000 to 87.975 MHz band. Twentychannels may be preset. Audio signals are appliedthrough the No.2 position of the transmitter-interphone selector switches and through the No.2receiver audio switches on the pilot’s and copilot’saudio control panels (fig. 3-1). Complete provisionsonly are installed for a TSEC/KY-28 voice securitydevice. Circuits are protected by a lo-ampere VHFAM-FM circuit breaker on the overhead circuitbreaker panel (fig. 2-26). Figure 3-5 illustrates theVHF AM-FM control panel. The associated antennais shown in figure 2-1.

3-9

TM 55-1510-221-10

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

3-10

TM 55-1510-221-10


(1.) 10 MHz selector. Selects receiver-transmitter frequency in increments of 10 MHzfrom 30 to 150 MHz. Clockwise rotation increasesfrequency.

(2.) 10 MHz indicator. Indicates manuallyselected receiver-transmitter frequency in 10 MHzincrements from 30 to 150 MHz.

(3.) 1.0 MHz selector. Selects receiver-transmitter frequency in 1.0 MHz increments.Clockwise rotation increases frequency.

(4.) I.0 MHz indicator. Indicates manu-ally selected receiver-transmitter frequency in 1.0MHz increments.

(5.) 0.1 MHz indicator. Indicates manu-ally selected receiver-transmitter frequency in 0.1MHz increments.


(7.) 0.025 MHz indicator. Indicates manu-ally selected receiver-transmitter frequency in 0.025MHz increments,


(9.) Preset CHAN indicator. Indicatesselected preset channel.

(10.) Preset CHAN selector. Selects presetchannel from 1 to 20. Clockwise rotation increasesnumber selected.

(11.) Preset channels freq. list. Writing areato keep track of preset channels,

(12.) LOCKOUT FM-AM switch. Screw-driver adjustable three-position switch. Warningtone announces lockout.

(a.) Center. Selects AM or FM band.

(b.) AM. Shuts off AM band.

(c.) FM. Shuts off FM band.

(13.) FM SQUELCH control. Screwdriveradjustable potentiometer. Squelch fully overdrivenat full counterclockwise position. Clockwise rotationincreases input signal required to open squelch.

(14.) WB-NB-MEM LOAD switch. Three-position switch.

(a.) NB. Limits selectivity to narrow-band intermediate frequency.

(b.) WB. Limits selectivity to wide-band intermediate frequency of FM band.

( c . ) M E M L O A D . Momentaryswitch. If pressed, loads manually selected frequencyin preset channel memory.

(15.) AM SQUELCH control. Screwdriveradjustable potentiometer. Squelch fully overriddenat full counterclockwise position. Clockwise rotationincreases input signal required to open squelch.

(16.) Mode selector switch. Three-positionrotary switch.

(a.) OFF. Shuts off receiver-transmitter.

modes.(b.) TR. Selects transmit/receive

(c.) DF. Not operational.

(I 7.) SQ-DIS-TONE select switch. Three-position switch.

tion.

(a.) Center. Selects squelch function.

(b.) SQ-DIS. Shuts off squelch func-

( c . ) TONE. T r a n s m i t s t o n e o fapproximately 1000 Hz.

(18.) Frequency control/emergency selectswitch. Three-position switch.

selection.(a.) PRE. Enables preset channel

selection.(b.) MAN. Enables manual frequency

(c.) EMER-AM-FM. Selects a pre-stored guard channel.

(19.) VOL control. Clockwise rotationincreases volume.


(1.) Turn-on procedure: Mode selectorswitch (VHF AM-FM control panel, fig. 3-5) -TR.


1. Frequency control emergencyselector switch (fig. 3-5) -MAN orPRE, as desired.

2. Transmitter-interphone selectorswitch (audio control panel, fig.

3-11

TM 55-1510-221-10

3-1) - No.2 position, or radiomonitor control No.2 -ON.

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



1. Transmitter-interphone selectorswitch (audio control panel, fig.3-1) - No.2 position.


(4.) Shutdown procedure: Mode selectorswitch (fig. 3-5) - OFF.

d. VHF AM-FM Emergency Operation:

(1.) Emergency AM Mode:


2. Mode selector switch - TR.

3. Frequency control/emergencyselector switch - EMER AM.

NOTE

Selecting EMER AM or FM automaticallydisables secure speech function andenables normal voice communication.


(2.) Emergency FM Mode:


2. Mode selector switch - TR.

3. Frequency control/emergencyselector switch - EMER FM.


(3.) Shutdown Procedure: Shutdown modeselect switch - OFF.

3-12. VOICE SECURITY SYSTEM TSEC/KY-28(PROVISIONS ONLY)

NOTE

Voice security system TSEC/KY-58 maybe installed in lieu of voice security sys-tem TSEC/KY-28. Complete provisions

are provided to install either the KY-28or the KY-58 voice security system on theLH fwd avionics rack behind the pilot(fig. 2-2).

a. Description. The KY-28 voice security sys-tem provides secure (ciphered) two-way voice com-munications for the pilot and copilot in conjunctionwith the UHF and VHF/AM/FM command sets,and the backup VOW set. System circuits are pro-tected by the VHF, VHF/AM/FM, RADIO RELAY,and BU VOW circuit breakers on the overhead cir-cuit breaker panel (fig. 2-26). Figure 3-6 illustratesthe KY-28 voice security (CIPHONY) control indi-cator.

b. Controls/Indicators and Functions. Voicesecurity control/indicator (LH fwd avionics rackbehind the pilot) (fig. 2-2).

Off.(1.) POWER ON switch. Turns set on or

NOTE

The POWER ON switch must be in ONposition for FM liaison or secure missionoperations in either the plain or ciphermode.

(2.) POWER ON indicator. Illuminateswhen POWER ON switch is placed in ON (up) posi-tion.

(3.) PLAIN indicator. Illuminates whenPLAIN/CIPHER switch is in PLAIN position.

(4.) PLAIN/CIPHER switch.

(a.) PLAIN. Enables uncipheredcommunications on FM liaison set.

(b.) CIPHER. Enables ciphered com-munications on FM liaison set.

(5.) RE-X/REG switch.

(a.) RE-X. Enables re-transmissionof ciphered communications at a distant location.

(b.) REG. Enables normal cipher orplain communications.

(6.) ZEROIZE switch. Normally OFF.Place in ON position during emergency situations toneutralize and make inoperative the associatedcipher equipment.

(7.) CIPHER indicator. Illuminates whenPLAIN/CIPHER switch is in CIPHER position.

3-12

TM 55-1510-221-10

Figure 3-6. Voice Security Control Indicator (C-8157/ARC)

c. VHF/AM/FM Set and Voice Security Oper-ation.

(1.) Turn-on procedure: POWER ONswitch (Voice security panel, fig. 3-6) - ON.

NOTE

The POWER ON switch must be in ONposition, regardless of the mode of theoperation, whenever the voice security(CIPHONY) KY-28 is installed in the air-craft.

(2.) Receive operating procedure:

1. SQUELCH control (VHF/AM/FM panel) - As required.

2. Transmitter-interphone selector(audio panel, fig. 3-1) -#2 posi-tion. or Audio monitor control #2- ON.

3. Mode selector (VHF/AM/FMpanel) - TR.

4. Frequency selectors (VHF/AM/FM panel) - As required.

5. PLAIN/CIPHER switch (voicesecurity panel) - As required.

(3.) Transmit operating procedure(PLAIN):

1. Transmit/interphone selector(audio panel) - No. 2 position.

2. PLAIN/CIPHER switch (Voicesecurity panel) - PLAIN.


(4.) Transmit operating procedure (CI-PHER) :

1. Transmit/inter-phone selector(audio panel) - No. 2 position.

PLAIN/CIPHER switch (Voicesecurity panel) - CIPHER. Indica-tor will be on while switch is inCIPHER position.)

2.

3.

4.

RE-X/REG switch (Voice securitypanel) - As required. (Set RE-Xposition only if distant station isusing re-transmitting equipment.)

Microphone switch - Pressmomentarily (interrupted tone

3-13

TM 55-1510-221-10

from voice security unit shouldno longer be heard).

NOTE

No traffic will be passed if the interruptedtone is still heard after pressing andreleasing the microphone switch.

5. Microphone switch- Press (do nottalk). Wait until beep is heard,then speak into microphone.


1. Mode selector (VHF/AM/FMpanel) - OFF.

2. POWER ON switch (Voice secur-ity panel) - OFF.

3-13. VOICE SECURITY SYSTEM TSEC/KY-58(PROVISIONS ONLY).

a. Description. The TSEC/KY-58 voice secur-ity system provides secure (ciphered) two-way voicecommunications for the pilot and copilot in con-junction with the UHF and VHF AM-FM commandsets, and the voice order wire set. The control indi-cator is located in the forward avionics rack behindthe pilot. System circuits are protected by the VHF,VHF AM-FM and BU VOW circuit breakers on theoverhead circuit breaker panel (fig. 2-26).


off.(1.) POWER ON switch. Turns set on or

NOTE

The power switch must be in ON positionfor FM or secure mission operations ineither the plain or cipher mode.

(2.) POWER ON indicator. Illuminateswhen POWER ON switch is placed in ON (up) posi-tion.

(3.) PLAIN indicator. Illuminates whenPLAIN-CIPHER switch is in PLAIN position.

(4.) PLAIN-CIPHER. Selects uncipheredor ciphered communications on FM set.

(a.) PLAIN. Enables uncipheredcommunications on FM set.

(b.) CIPHER. Enables ciphered com-munications on FM set.

(5.) RE-X-REG. Two-position switch.

(a.) RE-X. Enables re-transmissionof ciphered communications at a distant location.

(b.) REG. Enables normal cipher orplain communications.

(6.) ZEROIZE switch. Normally OFF.Place in ON position during emergency situations toneutralize and make inoperative the associatedcipher equipment.

(7.) CIPHER indicator. Illuminates whenPLAIN-CIPHER switch is in CIPHER position.

c. VHF AM-FM Set and Voice Security Oper-ation.

(1.) Turn-on procedure: Power switch(voice security panel, fig. 3-6) - ON.

NOTE

The power switch must be in ON posi-tion, regardless of the mode of the opera-tion, whenever the voice security(CIPHONY) KY-58 is installed in the air-craft.

(2.) Receive operating procedure:

1.

2.

3.

4.

5.

Squelch control (VHF AM-FMpanel, fig. 3-5) - As required.

Transmitter-interphone selectorswitch (audio control panel, fig.3-1) - No.2 position, or radiomonitor control No.2 -ON.

Mode selector switch (VHFAM-FM control panel, fig. 3-5) -TR.

Frequency selectors (VHFAM-FM control panel) - Asrequired.

Plain-cipher switch (voice securitycontrol panel) - As required.

(3.) Transmitter operating procedure(PLAIN):


2. Plain-cipher switch (Voice secur-ity control panel) -PLAIN.


3-14

(4.) Transmitter operating procedure (CI-PHER):

1. Transmitter-interphone selector(audio control panel, fig. 3-1) -No.2 position.

The HF system has two methods of frequencyselection. The first method is called direct tuning(frequency agile). The second is a channelized opera-tion in which desired operating frequencies are pre-set, stored and referenced to a channel number.

2. Plain-cipher switch (voice securitypanel) - CIPHER. (Indicator willbe illuminated while switch is inCIPHER position.)

b. Controls/Indicators and Functions (HFControl Panel, jig. 3-7.

(1.)selected.

FREQ display. Displays frequency

3. RE-X-REG switch (voice securitypanel) - As required. (Set RE-Xposition only if distant station isusing re-transmitting equipment.)

4. Microphone switch - Pressmomentarily (interrupted tonefrom voice security unit shouldno longer be heard.)

(2.)LSB, AM, or

(3.)selected.

MODE display. Displays selectedUSB mode.

CHANNEL display. Displays channel

(4.) Light sensor. The light sensor is aphotocell which adjust brigntness of the display.

NOTE

(5.) MODE switch. The mode switch is amomentary pushbutton switch that selects LSB, AMor USB.

No traffic will be passed if the interruptedtone is still heard after pressing and

(6.) FREQ/CHAN switch. Transfers the

releasing the microphone switch.HF system from a direct frequency operation to achannelized form of operation.

5. Microphone switch - Press (do nottalk). Wait until beep is heard,then speak into microphone.

(7.) PGM (Program) recessed switch.Enables channelized data to be modified. The PGMmessage will be displayed whenever this switch isdepressed.


1. Mode selector switch (VHFAM-FM panel) - OFF.

2. Power switch (voice securitypanel) - OFF.

NOTE

3-14. HF COMMUNICATION SET (KHF-950).

a. Description. The HF command set (fig. 3-7)provides long-range voice communications withinthe frequency range of 2.0 to 29.99 MHz andemploys either standard amplitude modulation(AM), lower sideband (LSB), or upper sideband(USB) modulation. The distance range of the set isapproximately 2,500 miles and varies with atmo-spheric conditions. With the capability to preset andstore 99 frequencies for selection during flight, thesystem also allows for selection of other frequenciesmanually (direct tuning), or reprogramming of anypreset frequency. The system will automaticallymatch the antenna by keying the microphone. Powerto the system is routed through a 25 ampere circuitbreaker placarded HF PWR. The receiving portionof the system is protected by a 5 ampere circuitbreaker placarded HF REC. Both circuit breakersare located on the overhead circuit breaker panel.

The program mode must be used for set-ting or changing any of the 99 preset fre-quencies. Each of the 99 channels may bepreset to receive and transmit on separatefrequencies (semi-duplex), receive only, ortransmit and receive on the same fre-quency (simplex). The operating mode(LSB,USB or AM) must be the same forboth receive and transmit and can also bepreset.

(8.) Frequency/channel selector. Thisselector consists of two concentric knobs that con-trol the channel and frequency digits, plus the lateralposition of the cursor.

(a.) Frequency control. The outerknob becomes a cursor (flashing digit) control withthe FREQ/CHAN switch in the FREQ position. Theflashing digit is then increased/decreased with theinner knob.

(b.) Channel control. The outer knobis not functional when the FREQ/CHAN switch isin the CHAN position. The inner knob will provide

TM 55-1510-221-10

3-15

TM 55-1510-221-10

Figure 3-7. HF Control Panel (KCU-951)

channel control from 1 through 99, displayed at theright end of the display window.

(9.) STO (Store) recessed switch. Storesdisplayed data when programming preset channels.

(10.) OFF- VOLUME control. Appliespower to the unit and controls the audio outputlevel.

(11.) SQUELCH control. Provides variablesquelch threshold control.

(12.) CLARIFIER control. Provides 250 Hzof local oscillator adjustment.


(1.) Turn on procedure:

NOTE

It is presumed aircraft power is on andnormally used avionic circuit breakersremain depressed.

NOTE

Aircraft can be configured for either HFor VOW on position 4 of Audio controlpanel (fig. 3-1).

1. AVIONIC MASTER POWERswitch - ON.

2. OFF-VOLUME switch - Turnclockwise out of OFF position.Adjust volume as desired.

(2.) Frequency operation (Simplex only):OFF-VOLUME Switch - Turn clockwise out of OFFposition. Adjust volume as desired.

3-16

TM 55-1510-221-10

NOTE

ch digit of the frequency may be selectedinstead of dialing up or down to a fre-quency. The larger concentric knob isused to select the digit to be changed.This digit will flash when selected. Rota-tion of the knob moves the flashing cursorin the direction of rotation. After the digitto be changed is flashing, the smaller con-centric knob is used to select the numeraldesired. This process is repeated until thenew frequency has been selected. Theflashing cursor may then be stowed bymoving it to the extreme left or right ofthe display and then one more click. Thisstows the cursor behind the display untilneeded again. The cursor may be recalledby turning the concentric knob one clickleft or right.

(3.) Direct frequency tuning (Simplexonly).

1. FREQ/CHAN button out(FREQ).

2. Select desired mode (USB,LSB,orAM).

3. Select digit to be changed (outerknob), digit (cursor) will flash.

4. Select numerical value of digit(inner knob).

5. Stow cursor (or repeat procedurefor additional changes).

6. Tune antenna coupler (pressmicrophone button).

(4.) Channel Programming.

NOTE

There are three ways to set up a channel:Receive only, simplex, and semi-duplex.To gain access to channelized operation,depress FREQ/CHAN button. To utilizethe existing programmed channels (i.e. noprogramming required) use the small con-trol knob to select the desired channelnumber. Then momentarily key themicrophone to tune the antenna coupler.If channel programming is required, it isnecessary to activate the program modeas follows. With the FREQ/CHAN buttonin (CHAN), use a pencil or other pointedobject to push the PGM button in. Thebutton is an alternate action switch:push-on, push-off. The letters PGM will

appear in the lower part of the displaywindow and the system will remain in theprogram mode until the PGM button ispressed again.


1. Stow the cursor if a frequencydigit is flashing.

2. Select the channel to be preset.

3. Set the desired operating mode(LSB,USB,or AM).

4. Set the desired frequency. (Referto frequency tuning)

5. Push and release STO buttononce.

NOTE

"T" will flash in the display window, how-ever a receive only frequency is being set.The flashing "T" should be ignored.

NOTE

If another channel is to be set, the cursormust be stowed before a new channel canbe selected. Use the smaller concentricknob to select the channel and repeat thesteps for selecting a new frequency.

6. To return to an operating mode,push the PGM button.

(a.) Simplex operation: Setting achannel up for simplex operation (receive and trans-mit on the same

1.

2.

3.

4.

5.

6.

frequency).

FREQ/CHAN button in (cursorstowed).

PGM button in (PGM displayed).

Select channel to be preset.

Set mode (LSB,USB or AM).

Set desired frequency. (Refer tofrequency tuning)

Push and release STO buttontwice.

The first press of the STO button stores the fre-quency in the receive position and the second pressstores the same frequency in the transmit position.The second push also stores the cursor.

If another channel is to be reset, use the smallerconcentric knob to select the channel and repeat the

3-17

TM 55-1510-221-10

steps for selecting a new frequency. The cursor wasautomatically stowed. To return to one of the oper-ating modes, push the PGM button again.

(b.) Semi-duplex operation: Setting achannel for semi-duplex (transmit on one frequencyand receive on another).

1. Select channel to be preset.

2. Set desired frequency. (Refer tofrequency selection)

3. Set mode (LSB,USB, or AM).

4. Push STO button once.

5. Set transmit frequency.

6. Push STO button again.

If another channel is to be reset, use the smallerconcentric knob to select the channel and repeat thesteps.

7. To return to an operating mode,push the PGM button.

NOTE

The mode for each channel (LSB, USB orAM) is stored along with the frequency. Ifthe mode is changed, the system willreceive and transmit in the mode selectedfor transmit.

d. Shutdown. Off/Volume switch - OFF.

e. HF Command Set - Emergency operation.Not applicable.

3-15. EMERGENCY LOCATOR TRANSMITTER(ELT).

a. Description. An emergency locator trans-mitter is provided to assist in locating an aircraftand crew in the event an emergency landing isnecessitated. The output frequency is 121.5 and 243MHz simultaneously. Range is approximately line-of-sight. The transmitter unit has separate functioncontrol switches located on one end of the case. Inthe event the impact switch has been inadvertentlyactuated, the beacon can be reset by firmly pressingthe pushbutton RESET switch on the front of thecase. The RESET switch and a 3-position toggleswitch, placarded ARM, OFF and ON, also on thetransmitter case, may be actuated by inserting onefinger through a small, round, spring-loaded door onthe left side of the aft fuselage (fig. 3-8). The trans-mitter unit is accessible through a service panellocated on the bottom of the aft fuselage.


(1.) RESET switch. When pressed, resetstransmitter.

(2.) Function switch. Selects operatingmode of set.

(a.) ARM. Arms set to be actuatedby impact switch (normal mode).

(b.) OFF. Turns set off.

(c.) ON. Manually activates trans-mitter for test or emergency purposes.

3-18

TM 55-1510-221-10

Figure 3-8. Emergency Locator Transmitter (Narco 03716-0300)

3-19

TM 55-1510-221-10

3-16. DESCRIPTION.

The navigation equipment group provides thepilot and copilot with instrumentation required toestablish and maintain an accurate flight course andposition, and to make an approach on instrumentsunder Instrument Meteorological Conditions (IMC).The navigation configuration includes equipmentfor determining attitude, position, destination rangeand bearing, heading reference and groundspeed.

3-17. RADIO MAGNETIC INDICATORS (RMI).

a. Description. The pilot and copilot are eachprovided with identical radio magnetic indicators(RMI) (fig. 3-9) located on the instrument panel(fig. 2-29). Each unit serves as a navigational aid forthe respective user and, by means of individualsource select switches, will display aircraft magneticor directional gyro heading and VOR, TACAN, INSor ADF bearing information. The pilot’s RMI isprotected by the 1-ampere No.1 RMI circuit breakeron the overhead circuit breaker panel (fig. 2-26) andthe 3.0-ampere F13 fuse on the No.1 junction box.The copilot’s RMI is protected by the 1-ampereNo.2 RMI circuit breaker on the overhead circuit

Section III. NAVIGATION

breaker panel and the 3.0-ampere F9 fuse on theNo.1 junction box.


(1.) Pilot’s COMPASS No. 1 - No.2 switch.Selects desired source of magnetic heading informa-tion for display on pilot’s HSI and copilot’s RMI.

(a.) No.1. Selects compass systemNo.1 for display control.

(b.) No.2. Selects compass systemNo.2 for display control.

(2.) Copilot’s COMPASS No. l-No.2switch. Selects desired source of magnetic headinginformation for display on copilot’s HSI and pilot’sRMI.

(a.) No.1. Selects compass systemNo.1 for display.

(b.) No.2. Selects compass systemNo.2 for display.

(3.) RMI select switch. Selects which oftwo signals will be displayed on respective RMI sin-

Figure 3-9. Radio Magnetic Indicator (RMI) (332C-10)

3-20

TM 55-1510-221-10

gle- needle pointer, if single-needle switch is in theVOR-TACAN position.

(a.) VOR 1. Selects VOR 1 bearingsignals for display.

(b.) TACAN. Selects TACAN bearingsignal for display.

c. Indicators and Functions (RMI, fig. 3-9).

(1.) Double needle pointer. Indicates bear-ing selected by double needle switch.

(2.) Compass card. Indicates aircraft head-ing at top of dial.

(3.) Heading index. Reference point foraircraft heading.

(4.) Warning flag. Indicates loss of com-pass signal.

(5.) Double needle switch. Selects desiredsignal to be displayed by double needle pointer.

mation.(a.) ADF. Selects ADF bearing infor-

(b.) VOR. Selects VOR 2 bearinginformation.

(6.) Single needle pointer. Indicates bear-ing selected by single needle switch.

(7.) Single needle switch. Selects desiredsignal to be displayed on single needle pointer.

mation.(a.) INS. Selects INS bearing infor-

(b.) VOR-TACAN. Selects signal asdetermined by RMI select switch on instrumentpanel, either VOR 1 or TACAN.

3-18. HORIZONTAL SITUATION INDICATORS.

a. Description. The pilot and copilot have sep-arate HSI instruments on respective instrumentpanel sections (fig. 3-10 and 3-11). Each HSI com-bines displays to provide a map-like presentation ofthe aircraft position with respect to magnetic head-ing. Each indicator displays aircraft heading, coursedeviation, and glideslope data. The pilot’s HSIallows the desired course and heading to be input tothe autopilot. Course deviation data is supplied tothe HSI by the VOR 1 or VOR 2 systems, theTACAN, or the INS. Glideslope data is supplied bythe VOR 1 or VOR 2 systems. The HSI displayswarning flags when the VOR, TACAN, INS or glide-slope signals are lost or become unreliable.

Figure 3-10. Pilot’s Horizontal Situation Indicator (RD-650B)

3-21

TM 55-1510-221-10

b. Controls/Indicators and Functions (Pilot'sHSI, fig. 3-10).

(1.) Distance display. Provides digital dis-plays of DME/TACAN or INS waypoint distance.TACAN distance is displayed in 1/10 mile incre-ments. INS distance to waypoint is displayed inwhole mile increments. The display will show dasheswhen the distance input data is invalid or absent.

(2.) Rotating heading (azimuth) dial. Dis-plays gyro stabilized magnetic compass informationon a dial which rotates with the aircraft throughout360 degrees. The azimuth ring is graduated in 5degree increments.

(3.) Lubber line. Fixed heading markslocated at the fore (upper) and aft (lower) position.

(4.) HDG flag. Indicates loss of reliableheading information.

(5.) Heading bug. The notched orangeheading bug is positioned on the rotating headingdial by the heading knob, to select and display a pre-selected compass heading. Once set to the desiredheading, the heading bug maintains its position onthe heading dial. The difference between the bugand the fore (upper) lubber line index is the amountof heading select error applied to the flight directorcomputer. In the heading mode the ADI will displaythe proper bank commands to turn to and maintainthis selected heading.

(6.) Course display. Provides a digitalreadout of selected magnetic course.

(7.) Course pointer. The yellow coursepointer is positioned on the heading dial by theremote course knob, to a magnetic bearing that coin-cides with the selected course being flown. Thecourse pointer rotates with the heading dial to pro-vide a continuous readout of course error to thecomputer.

(8.) Bearing pointer. Indicates ADF orNAV relative bearing as selected by the bearingpointer source switch.

(9.) ADF annunciator. When illuminated,indicates ADF bearing information is being dis-played.

(10.) Bearing pointer source switch. Thebearing pointer source switch, located on the pilot’sHSI, provides for selecting between ADF or NAVbearing information as presented by the bearingpointer. Each push of the select switch alternatesselection of ADF or NAV. Upon power-up or fol-lowing long-term power interruption, NAV is dis-played.

3-22

(1 I.) NAV annunciator. When illuminated,indicates NAV bearing information is being dis-played.

(12.) NAV flag. Indicates loss of VOR,TACAN or INS information, or unreliable naviga-tion signal.

(13.) Compass synchronization annuncia-tor. The compass synchronization annunciator con-sists of a dot and X symbol display. When the com-pass system is in the slaved mode, the display willoscillate between the dot and X symbol, indicatingthe heading dial is synchronized with a gyro stabi-lized magnetic heading.

(14.) Course deviation dots. In VOR orTACAN operation, each dot represents 5 degreedeviation from the centerline (± 10 degrees). In ILSoperation, each dot represents 1 degree deviationfrom the centerline. In INS operation, each dot rep-resents 3.75 nautical miles deviation from center-line.

(15.) Aircraft symbol. The fixed miniatureaircraft symbol corresponds to the longitudinal axisof the aircraft and lubber line markings. The symbolshows aircraft position and heading with respect toa radial course and the rotating heading (azimuth)dial.

(16.) Course deviation bar. The course devi-ation bar represents the centerline of the selectedVOR, TACAN, INS or localizer course. The minia-ture aircraft symbol pictorially shows actual aircraftposition in relation to this selected course.

(I 7.) VERT flag. Covers glide slope pointerwhen not receiving glide slope information.

(18.) Glide slope pointer/scale. The glideslope pointer displays glide slope deviation. Thepointer is in view only when tuned to a localizer fre-quency. If the aircraft is below glide slope path, thepointer is displayed upward on the scale. Each doton the scale represents approximately 0.4 degree dis-placement.

(19.) To-from pointer. The to-from pointersaligned on the course pointer, are located 180degrees apart. One always points in the direction ofthe station, along the selected VOR radial ortowards the INS waypoint.

(20.) Navigation source annunciators. Fivedifferent annunciators display navigation datasources. They are: TAC for TACAN, GPS (notused), INS, NV2 for VOR 2, NV1 for VOR 1. WPTindicates arrival at INS waypoint.

(21.) Course knob (located on the pedestal).Positions the course pointer.

TM 55-1510-221-10

(22.) Heading knob (located on the pedes- PILOT SELECT annunciator will illumi-tal). Positions the heading bug to a preselected head- nate to notify the copilot that both pilotsing. have selected the same receiver.

(23.) Pilot’s COURSE INDICATOR selectorswitch (fig. 2-29). Selects desired source of data fordisplay on pilot’s HSI and input to autopilot flightcomputer.

1 system.

2 system.

(a.) VOR 1. Selects data from VOR

(b.) VOR 2. Selects data from VOR

(I.) Compass synchronization annuncia-tor. The compass synchronization annunciator con-sists of a dot and X symbol display. When the com-pass system is in the slaved mode, the display willoscillate between the dot and X symbol, indicatingthe heading dial is synchronized with a gyro stabi-lized magnetic heading.

(c.) TACAN. Selects data fromTACAN system.

(2.) VERT flag. Indicates that the infor-mation displayed by the glideslope pointer is invalidand should not be used.

(d.) INS. Selects data from INS.

c. Controls/Indicators and Functions (Copi-lot’s HSI. fig 3-11).

(3.) Rotating heading (azimuth) dial. Dis-plays gyro stabilized magnetic compass informationon a dial which rotates with the aircraft throughout360 degrees. The azimuth ring is graduated in 5degree increments.

NOTE

If both the pilot and copilot COURSEINDICATOR select switches are in thesame position, except INS, the pilot hassole control of course select functions.The copilot can only monitor deviationdisplays from the selected system. A

(4.) Azimuth marks. Fixed azimuth marksare at 45° bearings throughout 360 degrees of thecompass card for quick reference.

(5.) Course deviation bar. The course devi-ation bar represents the centerline of the selectedVOR, TACAN, INS or localizer course. The minia-ture aircraft symbol pictorially shows actual aircraftposition in relation to this selected course.

Figure 3-11. Copilot’s Horizontal Situation Indicator (RD-550)

3-23

TM 55-1510-221-10

(6.) Digital COURSE counter. Provides adigital readout of selected magnetic course.

(7.) HDG flag. Indicates loss of reliableheading information.

(8.) Lubber line marks. Fixed headingmarks located at the fore (upper) and aft (lower)position.

(9.) Course pointer. The yellow coursepointer is positioned on the heading dial by thecourse knob to select a magnetic bearing that coin-cides with the desired VOR or TACAN radial orINS or localizer course. The course pointer rotateswith the heading dial to provide a continuous read-out of course error to the computer.

(10.) DIST display. Provides digital displayof station distance.

(Il.) Heading bug. The notched orangeheading bug is positioned on the rotating headingdial by the heading knob, and displays preselectedcompass heading. The bug rotates with the headingdial.

(12.) Bearing pointer. The bearing pointerprovides magnetic bearing to a selected TACAN orVOR station or INS waypoint.

(13.) Glideslope pointer/scale. The glideslope pointer displays glide slope deviation. Thepointer is in view only when tuned to a localizer fre-quency. If the aircraft is below glide slope path, thepointer is displayed upward on the scale. Each doton the scale represents approximately 0.4 degree dis-placement.

(14.) NAV flag. Indicates that informationderived from the selected navigational source (VOR,TACAN or INS) is invalid and should not be used.

(15.) To-from pointers. The to-from point-ers aligned on the course pointer, are located 180degrees apart. One always points in the direction ofthe station, along the selected VOR or TACANradial or toward INS waypoint.

(16.) Heading knob. Positions the headingbug to a preselected compass heading.

(17.) Course deviation dots. In VOR,TACAN or INS operation, each dot represents a 5degree deviation from the centerline (° 10 degrees).In ILS operation, each dot represents 1 degree devi-ation from the centerline. In INS operation, eachdot represents a 3.75 nautical miles deviation fromcenterline.

(18.) Aircraft symbol. The fixed miniatureaircraft symbol corresponds to the longitudinal axis

3-24

of the aircraft and lubber line markings. The symbolshows aircraft position and heading with respect toa radio course and the rotating heading (azimuth)dial.

(19.) Course knob. Positions the courseindicator.

(20.) Copilot's COURSE INDICATORswitch fig. 2-29). Selects desired source of data fordisplay on copilot’s HSI.

1 system.(a.) VOR 1. Selects data from VOR

2 system.(b.) VOR 2. Selects data from VOR

(c.) TACAN. Selects data fromTACAN system.

(d.) INS. Selects data from INS.

3-19. PILOT’S ATTITUDE DIRECTOR INDICA-TOR.

a. Description. The pilot’s attitude directorindicator (ADI) (fig. 3-12) combines the attitudesphere display with computed steering informationto provide the commands required to intercept andmaintain a desired flight path. It also contains aneyelid display, expanded localizer, glide slope, radioaltitude display, rate-of-turn indicator, mode annun-ciators, go-around and decision height annunciators,and inclinometer. Any warning flag in view indi-cates that portion of information is unreliable.


(1.) Attitude sphere. Moves with respect tothe symbolic aircraft reference to display actualpitch and roll attitude. Pitch attitude marks are in 5degree increments on a blue and brown sphere.

(2.) Roll attitude index. Displays actualroll attitude through a movable index and fixedscale reference marks at 0, 10, 20, 30, 45, 60 and 90degrees.

(3.) GA (go-around) annunciator. Illumi-nates when go-around mode has been selected.

(4.) SPD annunciator. Illuminates whenairspeed is being held by the flight director, in theIAS mode.

(5.) ALT annunciator. Illuminates whenaltitude is being held by the flight director.

(6.) HDG annunciator. Illuminates whenheading is being held by the flight director, in theNAV ARM, BC ARM mode.

TM 55-1510-221-10

Figure 3-12. Pilot’s Attitude Director Indicator

(7.) NAV annunciator. Illuminates whennavigation is being controlled by the flight director,in the NAV CAP, VOR APR mode.

(8.) LOC annunciator. Illuminates when-ever the flight director is controlling a localizerapproach, in the NAV CAP mode.

(9.) APR annunciator. Illuminates when-ever the flight director is controlling a approach, inthe NAV CAP, VOR APR mode.

(10.) GS annunciator. Illuminates wheneverthe flight director is in GS CAP mode, and glideslope has been captured.

(11.) BC annunciator. Illuminates when-ever the flight director is in BC CAP mode, and hascaptured the back course approach heading.

(12.) VRT annunciator. Illuminates whenvertical speed is being held by the flight director, inthe VS mode.

(13.) DH annunciator. Illuminates whenaircraft descends below selected decision height asset on the radio altimeter indicator.

(14.) Eyelid display. Surrounds the attitudesphere and provides positive attitude identificationby means of a blue eyelid which always shows the

relative position of the sky, and a brown eyelidwhich always shows the relative position of theground. The eyelids maintain the proper ground-skyrelationship, regardless of sphere position.

(15.) Speed command display. The pointerindicates relative airspeed provided by the angle-of-attack/speed command system.

(16.) Flight director command cue. Displayscomputed commands to capture and maintain adesired flight path. Always fly the symbolic minia-ture aircraft to the flight director cue. The cue willbias from view should a failure occur in either thepitch or roll channel.

(17.) Radio altitude display. Radio altitudeis digital displayed. The range capability of the dis-play is from -20 to 2500 feet AGL. The display reso-lution between 200 and 2500 feet is in 10 foot incre-ments. The display resolution below 200 feet is 5feet. The display will be blank at altitudes over 2500feet AGL. Dashes are displayed whenever invalidradio altitude is being received.

(18.) DH SET control knob. Sets decisionheight from 0 to 990 feet. Decision height displaysin the DH window on lower left corner of ADI. Thebrightness of the digital radio altitude and decisionheight display is controlled by the dimming knob

3-25

TM 55-1510-221-10

which is concentric with the DH SET knob. Thedimming knob also dims the distance and coursedisplay on the pilot’s HSI, and the altitude alert dis-play.

(19.) Expanded localizer. Raw localizer dis-placement data from the navigation receiver (HSIdisplay) is amplified approximately 7 1/2 times topermit the expanded localizer pointer to be used asa sensitive reference indicator of the aircraft’s posi-tion, with respect to the center of the localizer. It isnormally used for assessment only, since the pointeris very sensitive and difficult to fly throughout theentire approach. During final approach, the pointerserves as an indicator of the Category II window.Full scale deflection of the expanded localizerpointer is equal to 1/4 degree of beam signal. Theexpanded localizer is displayed by the localizerpointer only when a valid localizer signal is avail-able.

(20.) Inclinometer. Gives the pilot a con-ventional display of aircraft slip or skid, and is usedas an aid to coordinated maneuvers.

(21.) Rate of turn. Rate of turn is displayedby the pointer at the bottom of the ADI. The marksat the extreme left and right sides of the scale repre-sent a standard rate turn.

(22.) Attitude (ATT) test switch. W h e ndepressed, the sphere will show an approximate atti-tude change of 20 degrees of right bank at 10degrees pitch-up. The ATT warning flag will appear.In addition, all mode annunciator lights except DHwill illuminate.

(23.) Radio altitude (RA) test switch. Press-ing the RA test button causes the following displayson the radio altitude readout: all digits display 8’sthen dashes, and then the preprogrammed test alti-tude as set in the radio altimeter R/T unit, until thetest button is released at which time the actual alti-tude is displayed. The DH display during the testdisplays all 8’s with the altitude display and thendisplays the current set altitude for the remainder ofthe test. RA test is inhibited as a function of APRCAP.

(24.) Decision height (DH) display. The dig-ital DH display, displays decision height range from0 to 990 feet in 10 foot increments. The decisionheight is set by the knob in the lower right corner ofthe ADI.

(25.) Symbolic miniature aircraft. Serves asa stationary symbol of the aircraft. Aircraft pitchand roll attitudes are displayed by the relationshipbetween the fixed miniature aircraft and the mov-able sphere. The symbolic aircraft is flown to alignthe command cue to the aircraft symbol in order to

3-26

satisfy the commands of the selected flight directormode.

(26.) Glide slope scale and pointer. Displaysaircraft deviation from glide slope beam center onlywhen tuned to a ILS frequency and a valid glideslope is present. The aircraft is below glide path ifpointer is displaced upward. The glide slope dot rep-resents approximately 0.4 degree deviation from thebeam centerline.

3-20. COPILOT’S GYRO HORIZON INDICATOR.

a. Description. The copilot’s gyro horizonindicator (fig. 3-13) is a flight aid which indicatesthe aircraft’s attitude. The attitude given is in rela-tionship to an artificial horizon. There are no frontpanel fuses or circuit breakers provided for the copi-lot’s gyro horizon indicator.

b. Indicators and Functions.

(1.) Bank angle scale. Indicates aircraftbank angle from zero to 90 degrees with marks at10, 20, 30, 45, 60, and 90 degrees.

(2.) Bank angle index. Reference indicat-ing zero-degree bank.

(3.) Bank angle pointer. Indicates aircraftbank angle.

(4.) Horizon line. Affixed to sphere,remains parallel to the earth’s horizon at all times.

(5.) G flag. Presence announces loss ofpower.

(6.) Sphere. Indicates orientation withearth’s axis at all times.

(7.) Inclinometer. Assists the copilot inmaking coordinated turns.

(8.) Miniature aircraft. Indicates attitudeof aircraft with respect to the earth’s horizon.

3-21. TURN AND SLIP INDICATORS.

a. Description. The pilot and copilot haveidentical turn and slip indicators (fig. 3-14) pro-tected by the circuit breaker placarded TURN &SLIP on the overhead circuit breaker panel (fig.2-26).


(1.) Two-minute turn marks. Fixed mark-ers indicate two-minute turn rate when covered byturn rate indicator.

TM 55-1510-221-10

Figure 3-13. Copilot’s Gyro Horizon Indicator (GH-14B)

Figure 3-14. Turn and Slip Indicator (329T-1)

3-27

TM 55-1510-221-10

(2.) Turn rate indicator. Deflects to indi-cate rate of turn.

(3.) Inclinometer. Indicates lateral acceler-ation (side slip) of aircraft.

3-22. GYROMAGNETIC COMPASS SYSTEMS.

a. Description. Two identical compass sys-tems provide accurate directional information forthe aircraft at all latitudes of the earth. As a headingreference, two modes of operation are used: direc-tional gyro (FREE) mode, or slaved (SLAVE) mode.In polar regions of the earth where magnetic headingreferences are not reliable, the system is operated inthe FREE mode. In this mode, the system furnishesan inertial heading reference, with latitude correc-tions introduced manually. In areas where magneticheading references are reliable, the system is oper-ated in the SLAVE mode. In this mode, the direc-tional gyro is slaved to the magnetic flux detector,which supplies long-term magnetic reference to cor-rect the apparent drift of the gyro. Magnetic headinginformation from both systems is applied to variousaircraft systems through pilot and copilot COM-PASS No.1 - No.2 switches. There are no circuitbreakers for the gyromagnetic compass systems. Thecircuits are protected by the 2-ampere F2 and F6fuses on the No. 1 junction box.

b. Vertical Gyro A vertical gyro provides line-of-sight stabilization to the weather radar and rolland pitch information to the autopilot. A FASTERECT switch at the top of the pilot’s instrumentpanel (figure 2-29) provides a means for fast erec-tion of the gyros. Pressing and holding the FASTERECT switch will erect the gyro to within 1.0° ofpitch and roll within 60 seconds of power applica-tion, and erect to within 0.5° within 2 minutes. Nor-mal operation of the vertical gyro system will notrequire use of the fast erect switch. The circuit isprotected by the 3-ampere F22 fuse in the No. 1junction box.

c. Controls and Functions.

(1.) Pilot’s COMPASS No.1-No.2 switch.Selects desired source for magnetic heading informa-tion to display on

(a.)No. 1 for display.

(b.)No.2 for display.

pilot’s HSI and copilot’s RMI.

No. 1. Selects compass system

(3.) Toing:

1.

No.2. Selects compass system 2.

(2.) Copilot’s COMPASS No. I-No.2switch. Selects desired source for magnetic headinginformation to display on copilot’s HSI and pilot’sRMI and INS.

3-28

(a.) No.1. Selects compass systemNo. 1 for display.

(b.) No.2. Selects compass systemNo. 2 for display.

(3.) GYRO SLAVE-FREE switch. Selectssystem mode of operation.

(a.) SLAVE. Selects slaved mode.Compass flux valve connects to azimuth card.

(b.) FREE. Selects free mode. Fluxvalve is not connected to azimuth card.

(4.) INCREASE-DECREASE switch. Pro-vides manual fast synchronization of the system.

(a.) INCREASE. Causes gyro head-ing output to increase (move in clockwise direction).

(b.) DECREASE. Causes gyro head-ing output to decrease (move in counter-clockwisedirection).

d. Normal Operation.

(1.) Alignment procedure:

1. Gyro compass slave-free switch -SLAVE.

2.

(2.) To

1.

2.

3.

Gyrocompass increase-decreaseswitch - Hold switch momentarilyin the direction desired, and thenrelease. This will place system infast erect mode. The gyro willthen erect at approximately 30degrees per minute. While in thefast erect mode, the HEADINGflag (HSI) will be in view. Whenthe HEADING flag retracts fromview, the heading displayed willbe the magnetic heading.

determine magnetic heading:

Gyrocompass slave-free switch -SLAVE.

RMI rotating heading dial (com-pass card) - Read heading.

determine directional gyro head-

Gyrocompass slave-free switch -FREE.

Gyrocompass increase-decreaseswitch - Hold until the RMI com-pass card aligns with the magneticheading, then release.

Read heading. The heading willagree with the appropriate HSI.

e. Shutdown Procedure. Both compass sys-tems are shut down when the INVERTER No.1 orINVERTER No.2 switch is turned off. (If eitherinverter is on, both compass sets will be energized.)

3-23. ALTITUDE SELECT CONTROLLER

The Altitude Select Controller (fig. 3-15) pro-vides a means for setting the desired altitude refer-ence for the altitude alerting and altitude preselectsystem.

(1.) Altitude Alert. As the aircraft reachesa point 1000 feet from the selected altitude, a signalis generated to light the warning light on the altime-ter. This light remains on until the aircraft is 250feet from the selected altitude. If the aircraft nowdeviates by 250 feet or more from the selected alti-tude, the light is again energized. The light remainson until the aircraft returns to within 250 feet ordeviates more than 1000 feet from the selected alti-tude.

(2.) Altitude preselect The altitude isselected by turning the selector knob until the alti-tude display reads the desired value. No furtheraction is taken on the controller. To initiate altitudepreselect, the ALTSEL button is selected on theflight director controller. The pilot must initiate a

TM 55-1510-221-10

maneuver to fly toward the preselected altitude. Anyof the following pitch modes may be engaged: PitchHold, Airspeed Hold or Vertical Speed Hold. Uponinitiation of altitude preselect capture, the previ-ously selected pitch mode is automatically reset.

3-24. RADIO ALTIMETER INDICATOR.

Description. The radio altimeter indicator(fig. 3-16) displays radio altitude information from2500 feet to touchdown with an expanded linearscale under 500 feet.


(1.) DH annunciator. Light illuminates toalert that aircraft is at or below selected DH.

(2.) Decision height bug. Manually set byknob to establish DH.

(3.) Failure warning flag (not shown).When visible, indicates that system information maybe unreliable.

(4.) Altitude pointer. Points to dial readingfor current radio altitude from 0 to 2500 feet.

(5.) Decision height set knob. Used tomanually set DH.

Figure 3-15. Altitude Select Controller (AL-800).

3-29

TM 55-1510-221-10

(6.) TEST pushbutton. Pressed to checkindicator R/T unit and flag operation.

Operating the test button causes the flag to comeinto view and altitude pointer to indicate approxi-mately 100 feet. Release of button causes pointer toreturn to existing altitude and flag to retract.

3-25. VOR/LOC NAVIGATION SYSTEM.

a. Description. The aircraft is equipped withtwo VOR systems, controlled by a dual NAV 1 -NAV 2 control panel located on the pedestal (fig.2-7). Either VOR can direct input signals to the atti-tude director indicator. Controls are shown on figure3-1. Each VOR system includes independentreceiver units for VOR/LOC and glideslope (GS).Each VOR receiver provides a VOR input to arespective RMI, HSI, and the flight director com-puter. Each glideslope receiver sends GS flag andpointer deviation information to the HSI and flightdirector computer. VOR/LOC indicators may beused for navigation during manual control of the air-craft, or the autopilot may be coupled to the VORsystem, accepting VOR inputs to the autopilot com-puter. The pilot’s unit (VOR 1) is a navigation radiosystem which receives and interprets VHF omni-directional radio range (VOR) and localizer (LOC)signals, glideslope signals (GS), and marker beacon

3-30

Figure 3-16. Radio Altimeter Indicator (RA-315)

signals. It has a maximum range of 120 nauticalmiles line-of-sight. The system operates in a VOR/LOC frequency range of 108.00 to 117.95 mega-hertz, in a glideslope frequency range of 329.15 to335.00 megahertz, and at a marker beacon fre-quency of 75 megahertz. VOR 2 is similar to VOR1 except VOR 2 cannot receive or interpret markerbeacon signals. Each VOR system provides coursedeviation and glide path data, which can beswitched either to the copilot’s HSI or to the autopi-lot flight computer and pilot’s HSI, or both. Theaudio outputs of VOR 1 and VOR 2 systems aresupplied to the NAV control on the audio controlpanels. VOR 1 bearing data is supplied to the single-needle pointer on both Radio Magnetic Indicators.VOR 2 bearing data is supplied to the double-needlepointer on both Radio Magnetic Indicators. VOR 1uses a marker beacon antenna located on the under-side of the forward fuselage (fig. 2-1). VOR 1 andVOR 2 both use the same glideslope antenna,located inside the radome. Both VOR’s are pro-tected by separate 2-ampere circuit breakers, locatedrespectively on the number 1 and number 2 avionicsbus. The circuit breakers are placarded VOR No.1and VOR No.2 are located on the overhead circuitbreaker panel (fig. 2-26).

TM 55-1510-221-10

Figure 3-17. NAV 1 - NAV 2 Control Panel (VIR-30AGM, VIR-30AG

b. Controls/Indicators and Functions (NAV 1 (2.) Frequency control.Control Panel, fig. 3-17). frequency of VOR 2 receiver.

(1.) Frequency indicator. Displays selectedfrequency of VOR 1 receiver.

(3.) NAV-TEST switch.

(2.) Frequency control. Selects operatingfrequency of VOR 1 receiver.

VOR 2 navigation systems. If the system is function-ing properly the following indications will be pres-ented:

(3.) NAV-TEST pushbutton. Activates testof VOR 1 navigation system. If the system is func-tioning properly, the following indications are pres-ented:

(4.) OFF/VOL control. Activates VOR 2receiver. Permits monitoring VOR 2 audio andadjusts volume of signals received.

(a.) RMI. Single needle indicates 0°.(a.) RMI. Single needle indicates 0°.

(b.) HSI. Indicates lateral deviationto the right and glideslope deviation down, if flag istuned to ILS frequency. It tuned to NAV frequency,indicates 0° and the G/S flag is in view,

(b.) HSI. Indicates lateral deviationto the right and glideslope deviation down, if NAVis tuned to ILS frequency. If tuned to NAV fre-quency, indicates 0° and G/S flag is in view.

(4.) VOL-OFF control. Activates VOR 1receiver. Permits monitoring VOR 1 audio andadjusts volume of signals received.

c. Controls/Indicators and Functions (NAV 2Control Panel, Fig. 3-17).

(1.) Frequency indicator. Displays selectedfrequency of VOR 2 receiver.

d. Controls and Functions, Instrument Panel.

(1.) Pilots COURSE INDICATOR switch.Selects VOR receiver to control pilot’s HSI.

HSI.(a.) VOR 1. VOR 1 controls pilot’s

HSI.(b.) VOR 2. VOR 2 controls pilot’s

Selects operating

Activates test of

3-31

TM 55-1510-221-10

(2.) Copilot's COURSE INDICATORswitch. Selects VOR receiver to control copilot’sHSI.

HSI.(a.) VOR 1. VOR 1 controls copilot’s

HSI.(b.) VOR 2. VOR 2 controls copilot’s

(3.) NAV-A switch (audio control panel,fig. 3-1). Applies VOR audio to respective headsets.

(4.) MKR BCN HI-LO (marker beaconaudio control panel, fig. 3-2). Controls sensitivity ofmarker beacon receiver.

e. VOR Operation.

(1.) Turn-on procedure:

1. Insure that aircraft DC and ACpower is on.

2. Avionics master power switch(overhead control panel, fig. 2-12)- ON.

3. Frequency controls (VOR controlpanel) - Set for both receivers.

4. Volume controls (VOR controlpanel - Turn clockwise to activatesets and adjust volume.

5. NAV A, then NAV B audioswitches (audio control panel, fig.3-1) - ON. Confirm proper signal,then OFF.

6. RMI and HSI - Confirm properindications.

(2.) Normal Operation.

1.

2.

3.

4.

Pilot/copilot course indicatorswitches (instrument panel) -Select VOR source.

To determine course to station onpilot’s HSI: TO-FROM pointerreads TO (up) position.

To determine bearing from sta-tion on pilot’s HSI: TO-FROMpointer reads FROM position(down).

To determine course to station onRMI: Select VOR, verify singleneedle points course to station.

(3.) Localizer (LOC) operation:

1. VOR frequency knob (NAVpanel) - Select frequency.

2. Pilot’s, copilot’s COURSE INDI-CATOR switches (instrumentpanel) - Select VOR source.

3. Check glideslope flags unmasked.

(4.) VOR receiver operation:

1. VOR frequency control knob(VOR control panel, fig. 3-17) -Select desired frequency.

2. Volume control knob (VOR con-trol panel, fig. 3-17) - Full on.

3. NAV A, audio switch - ON.Adjust audio.

(5.) Shutdown procedure: Volume control(VOR control panel, fig. 3-17) - OFF.

3-26. MARKER BEACON RECEIVER.

a. Description. The marker beacon receiver islocated inside the No.1 VOR receiver. The markerbeacon receiver obtains power through the VORreceiver. The marker beacon provides visual andaural indication of the aircraft’s position over a 75MHz marker beacon ground transmitter. Uponentering the range of marker beacon signals, blue,amber, or white annunciator lights will illuminate,and corresponding aural signals will indicate aircraftpassage over the (0) outer, (M) middle, (A) inner orairway marker beacons. Range is vertical to 50,000feet. Volume, and sensitivity controls are located onthe marker beacon audio control panel (fig. 3-1).


(1.) Marker beacon sensitivity switch(marker beacon audio control panel, fig. 3-2). Con-trols sensitivity of marker beacon receiver.

(a.) HI position. Enables high sensi-tivity operation of marker beacon receiver.

(b.) LO position. Enables low sensi-tivity operation of marker beacon receiver.

(2.) "0" indicator. Illuminates when air-craft passes over an outer marker beacon.

(3.) "M" indicator. Illuminates when air-craft passes over a middle marker beacon.

(4.) "A" indicator. Illuminates when air-craft passes over an inner or airway marker beacon.

c. Marker Beacon Operation.

1. Marker beacon volume control (markerbeacon audio control panel, fig. 3-2) -As required.

3-32

TM 55-1510-221-10

2. Marker beacon HI-LO switch (markerbeacon audio control panel, fig. 3-2) -As required.

3. Marker beacon indicator lights(instrument panel, fig. 2-29) - Confirmbeacon indication.

3-27. AUTOMATIC DIRECTION FINDER (DF-203).

a. Description. The Automatic DirectionFinder (ADF) (fig. 3-18), is a radio navigation sys-tem which provides a visual indication of aircraftbearing, relative to a selected ground radio station.It may also be used to home to a selected station,find aircraft position, or monitor conventionalmedium frequency AM radio transmissions. Thesystem is designed to provide reliable reception of a400-watt radio station at a range of 65 nauticalmiles throughout a 360-degree turn of the aircraft. Itoperates in a frequency range of 190 to 1750 kilo-hertz. Bearing indications are displayed visually onthe RMI's and aural signals are applied to the audiocontrol panels. The ADF system consists of areceiver, located on the forward side of the aft cabinbulkhead inside the pressure vessel; a control unit,located on the pedestal extension; a non-directionalsense antenna, installed in the aircraft dorsal tin; adirectional loop antenna, located on the underside

of the fuselage; and a quadrangle error corrector,installed on the loop antenna (to compensate for thedeflection of arriving radio signals by the wings andfuselage of the aircraft). The system is protected bya l-ampere ADF, a 5-ampere RADIO RELAY, anda 35-ampere AVIONICS BUS FEEDER No.2 circuitbreaker located on the overhead circuit breakerpanel (fig. 2-26).

The only warning that the crewmemberhas for an unreliable ADF signal or lossof the ADF receiver is the loss of theADF audio signal in the crewmember’sheadset. This signal should be monitoredduring all phases of the approach in IMCconditions.

NOTE

Keying the HF radio set while operatingthe ADF radio set will cause a momentar-ily unreliable ADF signal.

Figure 3-18. ADF Control Panel (DF-203)

3-33

TM 55-1510-221-10

b. Controls and Functions (ADF ControlPanel, fig. 3-18).

(I.) L-LOOP-R control. Operative onlywhen the function switch is in the LOOP or ADFposition. Center position removes rotation signalsfrom the loop antenna and the ADF pointer on theRMl's. L (left) or R (right) of center applies rotationsignals to loop antenna and ADF pointer on RMI'sfor 360-degree rotation left or right. Center positionholds antenna position.

(2.) BFO-OFF. At BFO (on) setting, per-mits line tuning with Beat Frequency Oscillator(BFO). Also provides audio tone when receivingunmodulated CW. OFF position turns BFO off.

(3.) Tuning meter. Indicates relativestrength of received signals.

(4.) TUNE control. Selects operating fre-quency.

(5.) Range switch. Selects operating fre-quency band.

(6.) FREQUENCY indicator. Indicatesselected frequency.

(7.) Mode selector. Selects operatingmode.


(b.) ADF. Permits automatic direc-tion finding or homing operation.

(c.) ANT. Permits reception usingsense antenna.

(d.) LOOP. Permits audio-null hom-ing and manual direction finding operations.

(8.) GAIN control. Adjusts volume ofreceived signal.

c. Controls and FunctionsPanel, jig. 3-1).

(1.) NAV-B switch. O NADF audio to respective headset.

d. Controls and FunctionsAudio Panel, fig. 3-2).

(1.) FILTER V-OFF

(Audio Control

position applies

(Marker Beacon

switch. Selectswhether voice filter will be used with ADF audio.

(2.) FILTER R-OFF switch. Selectswhether range filter will be used with ADF audio.

3-34

NOTE

Range and voice filtering are cancelledwhen both FILTER R and FILTER V areON. Range and voice audio will be heard.

e. ADF Normal Operation.

(1.) To

1.

2.

3.

4.

5.

6.

7.

8.

9.

(2.) Toreceiving only.

1.

2.

3.

4.

(3.) Tofinding.

1.

2.

3.

4.

5.

6.

operate the set as ADF.

Mode selector - ADF.

BFO-OFF switch - BFO.

Range switch - Select frequencyrange.

Audio control panel (fig. 3-l)NAV B switch - On and adjusted.

Gain control - As required.

TUNE control - Rotate for maxi-mum reading on tuning meterand zero BFO beat.

BFO-OFF switch - OFF.

Double needle switch (RMI, fig.3-9) - ADF.

Double needle pointer (RMI, fig.3-9) - Read course to station.

operate set for sense antenna

Mode selector (ADF controlpanel, fig. 3-18 - ANT.

Range switch - Select operatingrange.

Tune control - Rotate for maxi-mum reading on tuning meter.

Gain control - As required.

operate set for aural-null direction

Mode selector (ADF controlpanel, fig. 3-18) - ANT.

BFO-OFF switch - BFO.

Range switch - Select operatingrange.

Tune control - Tune desired sta-tion.

Gain control - Adjust for mini-mum audio output.

Double needle switches (RMI, fig.3-9) - As required.

7.

8.

9.

10.

BFO-OFF switch - OFF.

Mode selector switch - LOOP.

Loop switch - L or R. Turn left orright until a null is reached (mini-mum sound in headsets).

Double needle on RMI (fig. 3-9) -Read course to station,

NOTE

The true null and direction to the radiostation may be indicated by either end ofthe single needle. This ambiguity must besolved to determine proper direction tothe station.

(4.) Shutdown procedure: Mode selectorswitch (ADF control panel, fig. 3-18) - OFF.

3-28. TACAN SYSTEMS.

a. Description. Two Tactical Air Navigation(TACAN) systems are provided. One is dedicated tothe INS and is used only for position updating, andprovides only DME information to the INS.Theother is used in conjunction with other avionics sys-tems, including the flight director system and theautopilot. For normal navigation TACAN is a radionavigation system which provides aircraft distanceand bearing information relative to a TACANground station. Both systems operate in the L bandfrequency range of 962 to 1213 MHz. Their range,though limited to line-of-sight, is designed to pro-vide reliable reception of a TACAN ground stationat a distance of 170 nautical miles at an aircraft alti-tude of 20,000 feet. The normal time required forthe systems to lock on to a selected ground stationsignal is three seconds. Both systems are protectedby a 2-ampere circuit breaker, placarded TACAN,located on the overhead circuit breaker panel (fig.2-26).

b. TACAN System (Non-INS dedicated). TheTACAN system (Non-INS dedicated) consists of arange unit (which includes the system transmitter)and a bearing unit, both located in the right noseavionics compartment; a distance indicator (fig.3-19) located on the instrument panel; a controlunit (fig. 3-20) located in the pedestal extension;and an antenna, located on the top of the fuselage.The TACAN system (non INS dedicated) operatesin conjunction with TACAN and VORTAC groundstations to provide distance, ground speed, time-to-station, and bearing-tostation data. It operates inthe L band frequency range on one of 252 preselec-ted frequencies, 126 X mode and 126 Y mode chan-nels. Course deviation to TACAN stations are dis-

TM 55-1510-221-10

played on the HSI’s. Distance, time-to-station, andground speed are displayed on the TACAN digitaldisplay (fig. 3-19). The ground speed and time-to-station are accurate only if the aircraft is flyingdirectly toward the ground station at a sufficient dis-tance that the slant range and ground range arenearly equal. The (Non-INS dedicated) TACAN sys-tem may be connected to and used with the autopi-lot system. When employed as the primary means ofnavigation, aircraft flight may be controlled manu-ally or by the autopilot. Indications of aircraft head-ing and bearing to ground stations are displayed onthe horizontal situation indicators. Magnetic bearingto a station is displayed by the RMI and pilot’s HSIbearing pointer. TACAN distance, ground speed,and time-to-station are all displayed on the TACANindicator located on the copilot’s instrument panel(fig. 2-29). TACAN distance is displayed on theHSI’s. The TACAN control panel (fig. 3-20) enablesselection of the TACAN frequency (channel) to beused, and provides for self-test of TACAN circuits.At the present time, most TACAN and VORTACstations are operated in the X mode. When Y modestations are operational, air navigation charts willdesignate the Y mode stations. The small (outer)control provides system power ON-OFF and stationidentifier tone, volume and control.

c. INS TACAN System. The INS TACAN sys-tem is coupled directly to INS circuits. It is dedi-cated only to updating the INS, is activated whenthe INS is operational, and is controlled only by theINS. The INS TACAN consists of a range unit anda distance indicator, both located on the INS equip-ment rack and both identical to counterparts in the(non-INS dedicated) TACAN and antenna, locatedon the underside of the fuselage (fig. 2-1). No con-trols are required or provided for the INS TACANsystem.

d. Controls/Indicators and Functions.

(1.) TACAN control panel.

1.

2.

3.

4.

5.

6.

TEST SWITCH. Activates systemself-test.

CHANNEL INDICATOR. Dis-plays selected TACAN channel.

X-Y SWITCH. Selects X or Ymode for TACAN channels.

VOL CONTROL.TACAN volume.

Adjusts

OUTER CHANNEL SELECTORKNOB. For manual selection oftens and hundreds part of channelnumber.

INNER CHANNEL SELECTORKNOB. For manual selection ofunits part of channel number.

3-35

TM 55-1510-221-10

Figure 3-19. TACAN Control Panel (AN/ARN-136)

Figure 3-20. TACAN Distance Indicator (SANS-706)

3-36

TM 55-1510-221-10

7. ON-OFF SWITCH. Activates ordeactivates system.

(2.) TACAN distance indicator.

1. NM INDICATOR. Displays slantrange distance in nautical milesfrom aircraft to selected TACANground station.

2. KT INDICATOR. Displaysground speed in knots.

3 . MIN INDICATOR. Displaystime to TACAN station in min-utes.

e. (Non-INS dedicated) TACAN Operation.

(I.) Turn-On procedure:

1. Power switch (TACAN controlpanel, fig. 3-20) - ON.


3. Course indicator switches (instru-ment panel, fig. 2-29) -TACAN.

4. Self-test procedure: Course knobon pilot’s HSI - Set to 180°, pressand hold TEST switch.

5. Pilot’s HSI course deviation bar -Centered, with course knob set to180 ±2 degrees, and TO-FROMindicator indicating TO.

6. RMI bearing pointers (fig. 3-9)and pilot’s HSI bearing pointer -Point to 180 degrees.

7. HSI course knob - Increase theselected course. The course devia-tion bar on a 180 ± 2 degrees TOindication and the bearing point-ers on each RMI indicator read180 degrees. Using the courseknob, increase the selected course,the course deviation bar willmove left. Decrease the selectedcourse, the course bar will center,then move to the right of center.Full scale deflection will be 10± 1°.

f. Normal Operating Procedure:

1. RMI single needle switch (fig. 3-9) -VOR-TACAN.

2. RMI selector switch (instrument panel, fig. 2-29) -TACAN.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

Course indicator selector switch (in-strument panel, fig. 2-29) - TACAN.

TACAN X-Y switch - As required.

TACAN channel selector knobs - Selectdesired channel.

Wait 5 seconds for signal acquisitionand lock-on.

If bearing lock-on is not obtained, per-form an inflight self-test to insure cor-rect operation of the system. Anytimea course indicator NAV or VOR LOCflag is in view, bearing, course devia-tion, and TO-FROM information maybe inaccurate and should be disre-garded.

Insure that audio station identificationsignal is correct for the ground stationselected.

RMI single-needle pointer and pilotHSI bearing pointer -Read bearing tostation.

HSI course control knob - Set desiredcourse.

HSI course deviation bar - Read devia-tion from selected course. Coursearrow will show wind correction anglewhen the course deviation bar is cen-tered and the aircraft is tracking theselected course.

TACAN indicator and pilot’s HSI -Read distance to station.

To determine course TO or courseFROM a TACAN station, rotate courseknob (pilot’s HSI) until course devia-tion bar is centered and the TO-FROMpointer reads TO or FROM.

To use TACAN during pilot-controlledflight, control aircraft by manual con-trols, responding to information dis-played on the flight director, RMI,HSI, TACAN, and other instruments.

To use TACAN with the autopilot,engage autopilot and monitor autopilotperformance on flight director, RMI,HSI and TACAN indicators. Verifyadherence to preset heading andcourse, and confirm the execution ofdisplayed steering commands.

3-37

TM 55-1510-221-10

(fig.

NOTE

The TACAN ground speed reading will beaccurate only when the aircraft is on acourse directly to or from the TACANstation.

g. Shutdown procedure: TACAN power switch3-20) - OFF.

3-29. AUTOMATIC FLIGHT CONTROL SYSTEM.

The RC-12H aircraft is certified withwingtip pods installed. Should the podsbe removed, the autopilot system must bereplaced with a standard C-12D autopilot.Effected wiring must also be changed.

a. Description. The Automatic Flight ControlSystem is a completely integrated autopilot/flightdirector/air data system which has a full comple-ment of horizontal and vertical flight guidancemodes. These include all radio guidance modes, andair data oriented vertical modes.

When engaged and coupled to the flight director(FD) commands, the system will control the aircraftusing the same commands displayed on the attitudedirector indicator. When engaged and uncoupledfrom the flight director commands, manual pitchand roll commands may be inserted using the pitchwheel and turn knob.

When the autopilot is coupled, the flight directorinstruments act as a means to monitor the perfor-mance of the autopilot. When the autopilot is notengaged, the same modes of operation are availablefor flight director only. The pilot maneuvers the air-craft to satisfy the Flight Director commands, asdoes the autopilot when it is engaged.

b. Air Data Computer. A digital air data com-puter located in the forward avionics compartmentprovides the altitude information for the pilot’saltimeter indicator, altitude alerter, and transpon-der. The computer also provides altitude and air-speed hold function data to the flight control com-puters. The air data computer receives 28 VDCpower through, and is protected by, a 2 amp circuitbreaker placarded AIR DATA - ENCDR located inthe AVIONICS section of the overhead circuitbreaker panel. All air data computer functions areautomatic in nature and require no flight crewaction.

c. Flight Director Mode Selector. The FlightDirector Mode Selector (fig.3-22), located on thepedestal, provides for selection of all modes (exceptgo-around which is initiated by a remote switchlocated on the left power lever for the flight director.The top row of split light annunciated pushbuttonscontains the lateral modes and the bottom row con-tains the vertical modes. The mode buttons will illu-minate when manually selected, or automaticallyselected through other modes.

The split light pushbutton annunciators, illumi-nate amber for armed conditions and green for cap-tured. When more than one lateral or vertical modeis selected, the flight director system automaticallyarms and captures the submode. Mode annuncia-tions are also presented on remote annunciatorblocks, located above the pilot’s Attitude DirectorIndicator (ADI), and on the pilot’s ADI. Operatingmodes and annunciation events of the Flight Direc-tor system are detailed in figure 3-21.

d. Controls/Indicators and Functions (FDMode Selector, jig. 3-22):

(1 . ) Heading Mode Switch (HDG) .Engages heading mode. Commands aircraft toacquire the heading indicated by a heading markeron the pilot’s HSI.

(2.) Navigation Mode Switch (NAV).Engages the navigation mode selected.

(3.) VOR Approach Mode Switch (APR).Engages approach mode. Commands aircraft tointercept and track ILS inbound course.

(4.) Back Course Mode Switch (BC).Engages backcourse mode. Commands aircraft tointercept back course ILS.

(5.) VNAV Mode Switch. Not used.

(6.) Standby Mode Switch (SBY). Engagesstandby mode.

(7.) Indicated Airspeed Hold Mode Switch(IAS). Engages indicated airspeed hold mode.

(8.) Vertical Speed Hold Mode Switch(VS). Engages vertical speed hold mode.

(9.) Altitude Preselect Mode Switch (ALT-SEL). Engages altitude preselect mode.

(10.) Altitude Hold Mode Switch (ALT).Engages altitude hold mode. Commands aircraft tomaintain pressure altitude.

e. Autopilot Modes of Operation.

(1.) Heading Select Mode (HDG). In theHDG mode the flight director computer provides

3-38

fig 3-12

TM 55-1510-221-10

Figure 3-21. Flight Director Modes and Annunciators

3-39

TM 55-1510-221-10

Figure 3-22. Flight Director Mode Selector (MS-500)

inputs to the command cue to command a turn tothe heading indicated by the heading bug on theHSI. The heading select signal is gain programmedas a function of airspeed. When HDG is selected, itoverrides the NAV, BC APR and VOR APR modes.In the event of a loss of valid signal from the gyroor compass, the command cue on the AD1 is biasedout of view. automatically select

(2.) Navigation Mode (NAV). The Naviga-tion Mode represents a family of modes for variousnavigation systems including VOR, Localizer,TACAN and INS.

(a.) VOR Mode. The VOR Mode isselected by selecting either VOR 1 or VOR 2 on theCourse Indicator select switch on the pilot’s instru-ment panel, and then depressing the NAV button onmode selector with the navigation receiver tuned toa VOR frequency. (A VOR indicator, NV1 or NV2on the pilot’s HSI, will illuminate. VOR NAV infor-mation will display on the pilot’s HSI and RMI).Prior to VOR capture, the command cue receives aheading select command as described above and theHDG mode switch is illuminated along with theNAV ARM annunciators. Upon VOR capture thesystem automatically: switches to the VOR mode;HDG and NAV ARM annunciators extinguish;NAV capture (NAV CAP) annunciators will illumi-

nate. At capture, a command is generated to captureand track the VOR beam. The course error signal isgain programmed as a function of airspeed. Cross-wind washout is included which maintains the air-craft on beam center in the presence of crosswind.The intercept angle is used in determining the cap-ture point to ensure smooth and comfortable perfor-mance during bracketing.

When passing over the station, an overstationsensor detects station passage removing the VORdeviation signal from the command until it is nolonger erratic. While over the station, coursechanges may be made by selecting a new course onthe HSI.

If the NAV receiver is not valid prior to the cap-ture point, the lateral beam sensor will not trip andthe system will remain in the HDG mode. After cap-ture, if the NAV receiver, compass data or verticalgyro go invalid, the ADI command cue will bias outof view. Also, the NAV CAP annunciators willextinguish if the NAV receiver becomes invalid.

(3.) Localizer Mode. The Localizer Modeis selected by depressing the NAV button on themode selector with the navigation receiver tuned toa LOC frequency. Mode selection and annunciationin the LOC mode is similar to the VOR mode. Thelocalizer deviation signal is gain programmed as afunction of radio altitude, time and airspeed. If theradio altimeter is invalid, gain programming is afunction of glide slope capture, time and airspeed.Other valid logic is the same as the VOR mode.

(4.) VOR Approach Mode. The VORApproach Mode is selected by depressing the APRbutton on the mode selector with the navigationreceiver tuned to a VOR frequency and less than 10DME miles from the station. The mode operatesidentically to the VOR mode with the gains opti-mized for a VOR approach.

(5.) Back Course Mode. The back coursemode is selected by pressing the BC button on themode selector. Back course operates the same as theLOC mode with the deviation and course signalsreversed to make a back course approach on thelocalizer. When BC is selected, and outside the lat-eral beam sensor trip point, BC ARM and HDGannunciators will illuminate. At the capture point,BC CAP will be annunciated with BC ARM andHDG annunciators extinguished. When BC isselected, the glideslope circuits are locked out.

(6.) Localizer Approach Mode (APR). Theapproach mode is used to make an ILS approach.Pressing the APR button with a ILS frequencytuned, arms both the NAV and APR modes to cap-ture the localizer and glide slope respectively. Noalternate NAV source can be selected. Operating

3-40

LOC mode is the same as described above except, ifthe radio altimeter is invalid in APR mode, gainprogramming is a function of glide slope capture,time, and airspeed.

With the APR mode armed, the pitch axis canbe in any one of the other pitch modes except go-around. When reaching the vertical beam sensor trippoint, the system automatically switches to the glideslope mode. The pitch mode and APR ARM annun-ciators extinguish and APR CAP annunciator illumi-nates on the controller. At capture, a command isgenerated to make a gradual interception of the glideslope beam. Capture can be made from above orbelow the beam. The glide slope gain is programmedas a function of radio altitude, time and airspeed.The APR CAP annunciator on the Mode Selectorwill extinguish if the GS receiver becomes invalidafter capture.

Glide slope capture is interlocked so that thelocalizer must be captured prior to glide slope cap-ture. If the glide slope receiver is not valid prior tocapture, the vertical beam sensor will not trip andthe system will remain in the pitch mode. After cap-ture, if the NAV receiver, GS receiver, compass dataor vertical gyro becomes invalid, ADI command cuewill bias out of view. If the radio altimeter is notvalid, the glide slope gain programming will be afunction of GS capture, time, airspeed, and the mid-dle marker.

(7.) Pitch Hold Mode. Whenever a rollmode is selected without a pitch mode, the ADIcommand cue will display a pitch attitude hold com-mand. The pitch attitude can be changed by pressingthe CWS button on the control wheel and maneu-vering the aircraft. The command cue will be syn-chronized to zero while the button is depressed.Upon release of the button, the pitch command willbe such as to maintain the new pitch attitude. In thepitch hold mode, the ADI command cue will bebiased out of view if the VG is not valid.

(8.) TACAN Mode. The TACAN mode isselected by selecting “TACAN“ on the course indica-tor selector switch, located on the pilot’s instrumentpanel. A TACAN annunciator, placarded TAC,located on the pilot’s HSI, will illuminate. TACANnavigation information will display on the pilot’sHSI and RMI.

NOTE

The TACAN receiver must be tuned to avalid TACAN frequency. TACAN func-tions are identical to VOR using TACANinformation rather then VOR signals. TheARM/CAP annunciation is the same as inVOR mode.

TM 55-1510-221-10

(9.) Altitude Hold Mode (ALT). The Alti-tude Hold Mode is selected by depressing the ALTbutton on the mode selector. When ALT mode isselected, it overrides the APR CAP, GA, IAS, VS,and ALTSEL CAP modes. In the ALT mode thepitch command is proportional to the altitude errorprovided by the air data computer. The altitudeerror signal is gain programmed as a function of air-speed. Depressing and holding the CWS buttonallows the pilot to maneuver the aircraft to a newAltitude Hold reference without disengaging themode. Once engaged in the Altitude Hold Mode, themode will be reset if the air data computer is notvalid and the ADI command cue will bias out ofview if the VG is not valid.

NOTE

If the Baro setting on the altimeter ischanged, a command is generated to flythe aircraft back to the original altitudereference.

(10.) Indicated Airspeed Hold Mode (IAS).The Indicated Airspeed Hold Mode is selected bydepressing the IAS button on the mode selector.When IAS is selected, it overrides the APR CAP,GA, ALT, VS, and ALTSEL CAP modes. In the IASmode the pitch command is proportional to airspeederror provided by the air data computer. Depressingand holding the CWS button allows the pilot tomaneuver the aircraft to a new Airspeed Hold refer-ence without disengaging the mode. Once engaged inthe IAS mode, the mode will be reset if the air datacomputer is not valid. The ADI command cue willbias out of view if the VG is not valid.

(11.) Vertical Speed Hold Mode (VS). TheVertical Speed Hold Mode is selected by depressingthe VS button on the mode selector. When VS isselected, it overrides the APR CAP, GA, ALT, ALT-SEL CAP, and IAS modes. In the VS mode, thepitch command is proportional to VS error providedby the air data computer. Depressing and holdingthe CWS button allows the pilot to maneuver theaircraft to a new Vertical Speed Hold reference with-out disengaging the mode. Once engaged in the VSmode, the mode will be reset if the air data com-puter is not valid. The ADI command cue will biasout of view if the VG is not valid.

(12.) Altitude Preselect Mode (ALTSEL).The Altitude Preselect Mode is selected by pressingthe ALTSEL button on the mode selector. Thedesired altitude is selected on the altitude preselectcontroller. Pitch hold, VS or IAS may be selected asa mode to fly to the selected altitude. When outsidethe altitude bracket trip point, the ALTSEL ARMannunciator along with the selected pitch mode is

3-41

TM 55-1510-221-10

illuminated on the mode selector. When reachingthe bracket altitude, the system automaticallyswitches to the ALTSEL CAP mode and the previ-ously selected pitch mode is cancelled. When thealtitude is reached, the ALTSEL CAP mode is auto-matically cancelled and the Flight Director switchesto the ALT hold mode. If the air data computer isnot valid, the altitude preselect mode cannot beselected. The ADI command cue will bias out ofview if the VG is not valid.

(13.) Standby Mode (SBY). The StandbyMode is selected by depressing the SBY button onthe mode selector. This resets all the other flightdirector modes and biases the ADI command cuefrom view. While depressed, SBY acts as a lamp testcausing all mode annunciators to illuminate and theflight director warning flag on the ADI to come inview. When the button is released, the mode annun-ciator lights extinguish and the flight director warn-ing flag retracts from view.

(14.) Go-Around Mode. The Go-AroundMode is selected by depressing the remote go-around switch. When selected all other modes arereset, and the remote go-around (GA) and yaw damp(YD ENG) annunciators will be illuminated. TheADI command cue receives a wings level command(zero command when roll is zero). The commandcue also receives the go-around command which is

a 7-degree visual pitch up attitude command. Select-ing GA disconnects the autopilot. However, the yawdamper remains on.

Once go-around is selected any roll mode can beselected. The wings level roll command will cancel.The go-around mode is cancelled by selectinganother pitch mode, or CWS.

f. Autopilot Controller.

(1.) Description. The autopilot controller(fig. 3-23), provides the means of engaging theautopilot and yaw damper as well as manually con-trolling the autopilot through the turn knob andpitch wheel. The autopilot system limits are shownin Table 3-1.

(2.) Controls/Indicators and Functions(Autopilot Controller, fig. 3-23):

1. PITCH WHEEL. Movement ofthe pitch wheel will cancel onlyALT HOLD, and ALTSEL CAP.With vertical modes of VS or IASselected on the mode selector,rotation of the pitch wheel willchange the respective displayedvertical mode reference. VS orIAS modes may be cancelled bypressing the mode button on the

Figure 3-23. Autopilot Con troller

3-42

2.

3.

4.

5.

6.

7.

mode selector. If VS or IAS arenot selected, the pitch wheelworks as described above. Thepitch wheel is always disabledduring a coupled glide slope.

BANK LIMIT PUSHBUTTON/ANNUNCIATOR SWITCH.Selection of the Bank Limit modeon the autopilot controller pro-vides a lower maximum bankangle while in the Heading Selectmode. LOW will illuminate onthe Bank Limit switch. The lowerbank limit is inhibited and LOWis extinguished during NAV modecaptures. If Heading Select isagain engaged, Bank Limit willagain be illuminated. PressingBank Limit when illuminated willreturn autopilot to normal banklimits.

SOFT RIDE PUSHBUTTON/ANNUNCIATOR SWITCH. Softride reduces autopilot gains whilestill maintaining stability in roughair. This mode may be used witha n y F l i g h t D i r e c t o r m o d eselected.

TURN KNOB. Rotation of theturn knob out of detent results ina roll command. The roll angle isproportional to and in the direc-tion of the turn knob rotation. theturn knob must be in detent (cen-ter position) before the autopilotcan be engaged. Rotation of theturn knob cancels any other previ-ously selected lateral mode.

YAW DAMP SWITCH. Whenthe autopilot is not engaged, theyaw damper may be utilized bydepressing the YD ENGAGEpushbutton.

AP ENGAGE PUSHBUTTON/ANNUNCIATOR SWITCH. TheAP ENGAGE switch is used toengage the autopilot. Engaging theautopilot automatically engagesthe yaw damper. The autopilotmay be engaged with the airplanein any reasonable attitude.

ELEV TRIM ANNUNCIATORS.The elevator trim annunciatorindicates UP or DN when asustained signal is being appliedto the elevator servo. The

TM 55-1510-221-10

annunciator should not beilluminated when engaging theautopilot.

(3.) Autopilot Disengagement. The autopi-lot is normally disengaged by momentarily depress-ing the control wheel AP DISC switch. The autopilotmay however be disengaged by any of the following:

1.

2.

3.

4.

5.

6.

7.

8.

Actuation of the control wheel APDISC button. Disengagement isconfirmed by 5 flashes of the APENG annunciator.

Pressing the respective verticalgyro FAST ERECT button.

Actuation of respective compassINCREASE-DECREASE switch.

Selection of Go-Around mode.Disengagement is confirmed bythe AP ENG annunciator flashing5 times and illumination of theGA and YD ENG annunciators.

Pulling the autopilot CONTROL& AFCS DIRECT circuit breaker.

Pressing the autopi lo t APENGAGE pushbutton.

Actuation of the manual electrictrim.

Any of the following malfunctionswill cause the autopilot to auto-matically disengage:

a. Vertical gyro failure.

b. Directional gyro failure.

c. Autopilot power or circuitfailure.

d. Torque limiter failure.

NOTE

Disengaging under any of the previousfour conditions will illuminate the APDISC annunciator and the MASTERWARNING light. Pressing the controlwheel AP DISC switch will extinguish theAP DISC annunciator.

(4.) Pitch Sync & Control Wheel Steering(CWS). The CWS push button located on the controlwheel (fig. 2-17) allows the pilot to manually changeaircraft attitude, altitude, vertical speed and/or air-speed without disengaging the autopilot. After com-pleting the manual maneuver, the CWS pushbuttonis released, and the autopilot will automatically

3-43

TM 55-1510-221-10

Table 3-1, Autopilot System Limits(Sheet 1 of 2)

MODE

YawDamper

CONTROL OR SENSOR PARAMETER VALUE

Yaw Control Engage Limit Unlimited

Autopilot Engage Limit Roil Up to ±90 degEngage Pitch Up to ±30 deg

Basic Turn Knob Roil Angie Limit ±30 degAutopilot Roil Rate Limit ±5 deg/sec

Pitch Wheel Pitch Angle Limit ±15 deg Pitch

Heading Hold Roil Angie Limit Less than 6 deg and noroll mode selected

HeadingSelect

Heading SELKnob on HSI orRemote slew knob onconsole

Roil Angle LimitRoil Rate Limit

±25 deg±3.5 deg/sec

VOR Course Knob,NAV Receiverand TACANReceiver

CaptureBeam Angieintercept(HDG SEL)Roil Angie LimitCourse Cut Limitat CaptureCapture Point

ON Course

up to ±90 deg

±25 deg±45 deg Course

Function of beam,beam rate, course error.Max trip point is 175micro-amps. Min trip pointis 30 microamps

Roil Angie LimitCrosswindCorrectionOver Station

±13 deg of RoilUp to ±45 degCourse Error

LOC or Course KnobAPR or andBC NAV Receiver

Course ChangeRoll Angie Limit

LOC CaptureBeam Intercept

Roil Angle LimitRoil Rate LimitCapture Point

NAV On-CourseRoll Angle LimitCrosswindCorrection LimitGain Programming

UP to ±90 deg±17 deg

Up to ±90 deg

±25 deg±5 deg/secFunction of Beam, BeamRate and Course Error.Max Trip Point is 175microamps. Min Trip pointis 60 micoramps.

±17 deg of roll±30 deg of courseerrorf (time and TAS) starts at1200 ft radio altitude, gainreduction = 1 to 5

3-44

TM 55-1510-221-10

Table 3-1, Autopilot System Limits(Sheet 2 of 2)

MODE

LOC orAPR orBC (cont)

CONTROL OR SENSOR PARAMETER VALUE

Glideslops CaptureGS Receiver and Beam Capture Function of beamAir Data and beam rate. TripComputer point is 30 microamps

Pitch Command ±10 degLimitGlideslope Damping Vertical velocityPitch Rate Limit f (TAS)Gain Programming f (time and TAS)

starts at 1200 ft radioaltitude, gain reduction =1 to .33f (Radio Alt) starts at 250ft. gain reduction = .33 at250 ft to 0 at 0 ft.

GA Control Switchon Throttles

Pitch CWS Switch onSync Wheel

ALT Hold Air DataComputer

VS Hold Air Data Computer

IAS Hold Air DataComputer

ALT Preselect Air Data Computer

Fixed Pitch-Up 7 deg Pitch UpCommand, Wings Level

Pitch Attitude ±20 deg maxCommand

ALT Hold Engage Range 0 to 50,000 ft

ALT Hold Engage Error ±20 ftPitch Limit ±20 degPitch Rate Limit f (TAS)

VERT Speed Engage 0 to ±6,000 ft/min.RangeVERT Speed Hold Engage ±30 ft/minErrorPitch Limit ±20 degPitch Rate Limit f (TAS)

IAS Engage Range 80 to 450 knotsIAS Hold Engage Error ±5 knotsPitch Limit ±20 degPitch Rate Limit f (TAS)

Preselect Capture Range 0 to 50,000 ftMaximum Vertical ±4,000 ft/minSpeed for CaptureMaximum Gravitational ±20gForce During CaptureManeuverPitch Limit ±20 degPitch Rate Limit f (TAS)

3-45

TM 55-1510-221-10

resynchronize to the vertical mode. Example: withIAS mode selected, the pilot may depress the CWSpushbutton and manually change airspeed. Oncetrimmed at the new airspeed the CWS pushbutton isreleased, and the autopilot will hold this airspeed. Ifa large pitch attitude change is made, the pilotshould trim the aircraft normally before releasingthe CWS button.

and by the 20-ampere circuit breaker on the front ofthe INS battery unit.

6. Controls/Indicators and Functions (INSMode Selector Unit, fig. 3-24).

N O T E

Either pilot’s CWS button will permitchanging of the autopilot regardless ofwhich pilot has control of the autopilot.However, use of the CWS will cancel theother pilot’s flight director GA mode.

(1.) READY NAV lamp. Illuminates toindicate INS high accuracy alignment has beenattained. If attained during ALIGN mode, lightremains illuminated until NAV mode is selected.Light illuminates momentarily during alignment, ifalignment is accomplished while in NAV mode.

(2.) BAT lamp. Illuminates to indicateINS shutdown due to low battery unit voltage.

(3.) Mode select knob. Controls INS acti-vation and selects operating modes.

3-30. INERTIAL NAVIGATION SYSTEM. (a.) OFF. Deactivates INS.

a. Description. The Inertial Navigation Sys-tem (INS) is a self-contained navigation and attitudereference system. It is aided by (but not dependentupon) data obtained from its own TACAN system,the GPS, the aircraft encoding altimeter and thegyromagnetic compass system. The position andattitude information computed by the INS is sup-plied to the automatic flight control system, weatherradar system, horizontal situation indicator, andradio magnetic indicators. In conjunction with otheraircraft equipment, the INS permits operation underInstrument meteorological Conditions (IMC). TheINS provides a visual display of present positiondata in Universal Transverse Mercator (UTM) coor-dinates or conventional geographic (latitude-longitude) coordinates during all phases of flight.When approaching the point selected for a legswitch, an ALERT light will illuminate informingthe pilot of an imminent automatic leg switch or theneed to manually insert course change data. TheINS may be manually updated for precise aircraftpresent position accuracy by flying over a referencepoint of known coordinates. The INS may beupdated automatically by the TACAN system or theGPS. Altitude information is automatically insertedinto the INS computer by an encoding altimeterwhenever the INS is operational.

(b.) STBY. Moving to STBY fromOFF mode: Starts fast warmup of system to operat-ing conditions; activates computer so informationmay be inserted; all INS controlled warning flagswill indicate warning. Moving to STBY from anyother mode: INS operates as if in attitude referencemode.

(c . ) ALIGN (ground use on ly ,parked). Moving to ALIGN from OFF mode: Level-ing starts after fast warmup heaters are off. Movingto ALIGN from STBY: Alignment starts if fast war-mup heaters are off. Moving to ALIGN from NAVmode: INS is not downmoded, but will allow auto-matic shutdown if overtemperature is detected.

(d.) NAV. Activates normal naviga-tion mode after automatic alignment is completed;must be selected before moving aircraft. Moving toNAV from STBY mode causes INS to automaticallysequence through STBY and ALIGN to NAV mode,if present position is inserted and aircraft is parked.NAV mode is used to shorten time in STBY and tobypass battery test, if stored heading is valid.

The Mode Selector Unit (MSU) (fig. 3-24) con-trols system activation and selects operating modes.

The Control Display Unit (CDU) (fig. 3-25) pro-vides controls and indicators for entering data intothe INS and displaying navigation and system statusinformation.

(e.) ATT. Activates attitude refer-ence mode. Used to provide only INS attitude sig-nals. Shuts down computer and CDU leaving onlyBAT and WARN lights operative. Once selected,INS alignment is lost.

c. Controls/Indicators and Functions, (INSControl Display Unit, fig. 3-25).

The INS system is protected by the 10-ampereINS AC POWER and the 5-ampere INS HTR ACPOWER circuit breakers on the mission AC/DCpower cabinet, by the 5-ampere INS CONTROL cir-cuit breaker on the overhead circuit breaker panel

(1.) HOLD key. Used with other CDUcontrols to stop present position display from chang-ing, in order to update position and to displayrecorded malfunction codes. Lights when pressedfirst time; goes out when pressed second time orwhen inserted data is accepted by computer. When

3-46

TM 55-1510-221-10

Figure 3-24. INS Mode Selector Unit

pressed second time, allows displays to resume showingchanging current present position.

(2.) ROLL LIM key. Allows selection of RollLimited steering mode. Press to select mode, key lights.Roll steering output is limited to 10 degrees. Presssecond time to exit mode, key light extinguishes. Rollsteering output returns to normal limit of 25 degrees.

(3.) Data display, left and right. Composedof lights which illuminate to display numbers, decimalpoints, degree symbols, left and right directions, andlatitude or longitude directions.

(4.) INSERT / ADVANCE / HI PREC key.Allows insertion of loaded data into computer.Enters displayed data into INS. When pressedbefore pressing any numerical key, alternates displayof normal and high precision data.

(5.) ALERT lamp. Illuminates amber toalert pilot 1.3 minutes before impending automaticcourse leg change. Extinguishes when switched tonew leg, if AUTO-MAN switch is set to AUTO.Flashes on and off when passing waypoint, ifAUTO-MAN switch is set to MAN. Light willextinguish if AUTO is selected or if a course change isinserted.

(6.) BAT lamp. Illuminates amber to indicateloss of 115 VAC power and INS operation on INSbattery power.

(7.) WARN /amp. Lights red to alert pilotINS self-test circuits have detected a system fault.Illumination may be caused by continuous orintermittent condition. Intermittent conditions lightWARN light until reset by TEST switch. If continuouscondition does not degrade attitude operation, light goesout when mode selector is set to ATT.

(8.) Keyboard. Consists of 10 keys forentering load 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 ofTACAN station data. MV/P and DISP (on keys “3” and“9”) are associated with loading and display of magneticvariation and magnetic heading. Pattern steeringparameters and DISP (on keys “5” and “9”) areassociated with loading and display of UTMcoefficients and waypoint move parameters.

(9.) CLEAR key. When pressed, illuminatesand erases data loaded into data displays or FROM-TOdisplay. Used to cancel erroneous data. After clearing,data loading can be resumed.

3-47

TM 55-1510-221-10

Figure 3-25. INS Control Display Unit

3-48 Change 4

number in all navigation computations. If INSERT/ADVANCE key is not pressed, computer will con-tinue using original numbers in all navigation com-putations; but distance/time information, based onnew leg, may be called up and read in data displays(in case of waypoints). When not in TACAN mixmode, TACAN station number is inserted to displayDIS/TIME information.

(10.) WYPT CHG key. When pressed,enables numbers in FROM-TO display to bechanged. If INSERT/ADVANCE key is pressed,computer will use navigation leg defined by new

(11.) AUTO-MAN TEST switch. This is adual purpose control. When the knob is pressedinward, the TEST switch function is engaged. Whenthe knob is rotated to either the AUTO or MAN set-ting, the control serves as a selector between thosemodes.

(a.) AUTO. Selects automatic legswitching mode. Computer switches from one leg tothe next whenever waypoint in TO side of theFROM-TO display is reached.

(b.) MAN. Selects manual leg switch-ing mode. Pilot must make waypoint changes manu-ally.

(c.) TEST. When pressed, performstest of INS lights and displays, remote lights andindicators controlled by 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. Con-tinued pressing of switch provides constant INS out-puts to drive cockpit displays in a predeterminedfashion.

NOTE

The INS can provide test signals to theHorizontal Situation Indicator (HSI) andconnected displays. Pressing TEST switchduring STBY, ALIGN, or NAV modeswill cause all digits on connected digitaldisplays to indicate “S’s“ and illuminatesthe HSI “WAYPOINT” and ALERTlights. Additional HSI test signals are pro-vided when INS is in ALIGN and thedata selector is at any position other thanDSRTK/STS. Under those conditions,pressing TEST switch causes the HSI toindicate heading, drift angle, and trackangle error - all at “0°“ or “30°“. At thesame time, cross track deviation is indi-cated at “3.75“ nautical miles (one dot)

TM 55-1510-221-10

right or left and INS-controlled HSI flagsare retracted from view.

Output test signal are also supplied to theautopilot when INS steering is selected. RotatingAUTO-MAN switch to AUTO and pressing TESTduring align furnishes a 15° left bank steering com-mand. A 15° right bank steering command is fur-nished when the AUTO/MAN switch is set to MAN.

(12.) FROM-TO display. Display numbersdefining waypoints of navigation leg being flown or,in the case of a flashing display, displays TACANstation being used. Waypoint numbers automaticallychange each time a waypoint is reached. Unlessflight plan changes during flight, the automatic legswitching sequence will always be 1, 2, 2 3, 3 4....89, 9 1, 1 2, etc.

(13.) Data selector. Selects data to be dis-played in data displays or entered into INS. Therotary selector has 10 positions. Five positions (L/LPOS, L/L WY PT, UTM POS, UTM WY PT andDSRTK/STS) also allow data to be loaded into datadisplay then inserted into computer memory.

(a.) TK/GS. Displays aircraft trackangle in left display and ground speed in right dis-play.

(b.) HDG/DA. Displays aircraft trueheading in left display and drift angle in right dis-play.

(c.) XTK/TKE. Displays cross trackdistance in left display and track angle in right dis-play.

(d.) L/L POS. Displays or enterspresent aircraft position latitude in left data displayand longitude in right data display. Both displaysindicate degrees and minutes to nearest tenth of aminute. This position also enables the insertion ofpresent position coordinates during alignment andpresent position updates.

(e.) L/L WY PT. Displays or enterswaypoint and TACAN station data, if used in con-junction with the waypoint/TACAN selector. Thisposition will also cause display of inertial presentposition data when the HOLD key is illuminated.

(f.) DIS/TIME. Displays distancefrom aircraft to TACAN station or any waypoint, orbetween any two waypoints in left display. Displaystime to TACAN station or any waypoint, or betweenany two waypoints, in right display.

(g.) WIND. Displays wind directionin left display and wind speed in right display, whentrue airspeed is greater than the air data systemlower limit (115 to 400 KIAS).

3-49

TM 55-1510-221-10

(h.) DSRTK/STS. Displays desiredtrack angle to nearest degree in the left data display,and INS system status in right data display.

(i.) UTM POS. Displays or entersaircraft position in Universal Transverse Mercator(UTM) coordinates, with northing data in kilome-ters in left display and easting data in kilometers inright display. The extra precision display showsmeters.

(j.) UTM WY PT. Displays or enterswaypoint and TACAN station data in UTM coordi-nates. Also enables loading and display of spheroidcoefficients if GRID and DISP keys are pressedsimultaneously.

(14.) Dim knob. Controls intensity of CDUkey lights and displays.

(15.) Waypoint/DME selector. Thumbwheelswitch, used to select waypoints for which data is tobe inserted or displayed. Waypoint station “0“ is fordisplay only and cannot be loaded with usable data.

d. INS - Normal Operating Procedures.

NOTE

The following data will be required priorto operating the INS: latitude and longi-tude (Geographical) or Universal Trans-verse Mercator (UTM) coordinates of air-c r a f t d u r i n g I N S a l i g n m e n t . T h i sinformation is necessary to program theINS computer during alignment proce-dure.

NOTE

When 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 stepto left as each subsequent digit is entered.The degree sign, decimal point, and colon(if applicable) will appear automatically.

(1.) Preflight Procedure.

Insure that cooling air is available to nav-igation unit before turning the INS on.

NOTE

Aircraft must be connected to a groundpower unit if INS alignment is performed

3-50

prior to engine starting. In this event, theengines must not be started until after theINS is placed in the NAV mode.

1. Applicable circuit breakers -Check depressed.

2. Mode selector switch (MSU, fig.3-23) - ALIGN. Confirm follow-ing:

a. FROM-TO display (CDU,fig. 3-24) indicates “1 2“.

b. INSERT/ADVANCE push-button light (CDU) illumi-nates.

c. BAT light (MSU, fig. 3-23)illuminates for approxi-mately 12 seconds at align-ment state “8”, then extin-guishes.

3. Dim knob (CDU) - Adjust foroptimum brightness of CDU dis-plays.

4. AUTO-MAN TEST switch(CDU) - AUTO.

5. Data selector (CDU) - L/L POS orUTM POS, as desired. Observecoordinates of last present posi-t ion pr ior to INS shutdownappear in data displays.

NOTE

Aircraft must not be towed or taxied dur-ing INS alignment. Movement of this typeduring alignment causes large navigationerrors. If aircraft is moved during align-ment, restart alignment by setting modeselector switch to STBY, then back toALIGN and reinserting present position.

NOTE

Passenger or cargo loading in the aircraftcould cause the type of motion whichaffects the accuracy of alignment. Anyactivity which causes the aircraft tochange attitude shall be avoided duringthe alignment period.

6. AUTO-MAN TEST switch(CDU) - Press and hold for test.Confirm following on CDU:

7. Left and right data displays indi-cate “88°88.8 N/S” and “88°88.8E/W“ respectively.

TM 55-1510-221-10

8.

9.

10.

11.

FROM-TO display indicates“8.8“.

The following pushbuttons andlights illuminate: ROLL LIM,HOLD, INSERT/ADVANCE,WYPT CHG, ALERT, BAT (onC D U a n d MSU), WARN,READY NAV and WYPT onpilot’s HSI.

AUTO-MAN TEST switch(CDU) - Release. Confirm datadisplays indicate coordinates incomputer memory.

If UTM coordinates are to beused, verify that appropriate gridcoefficients have been loaded.

(2.) Insert present position:

NOTE

Prior to pressing INSERT/ADVANCEpushbutton, any incorrectly loaded datacan be corrected by pressing the CLEARpushbutton and reloading correct data.

NOTE

While parked aircraft is undergoing align-ment, encoding altimeter will supply thefield elevation (aircraft pressure altitude)into INS.

NOTE

Once present position has been insertedand computer has advanced to alignmentstate “7“, present position cannot be rein-serted without downmoding to STBY andrestarting alignment.

NOTE

If longitude and latitude coordinates arebeing used, skip following step (a) -2- andproceed with step (b) -2-.

(a.) Insert UTM coordinates of air-craft present position:

1. Data selector - UTM POS.Observe that prior to initialload, INSERT/ADVANCEpushbutton light illuminates.

2. To load zone and eastingvalues - Press keys insequence, starting with “E“.Example: Zone 16, 425 kmEast = E16 425. Observethat zone and easting in kilo-meters appear in right datadisplay as keys are pressed.

3. INSERT/ADVANCE push-button - Press. Observepushbutton light remainsilluminated.

4. To load northing data -Press keys in sequence, start-ing with “N“ or “S“ to indi-cate north or south hemi-sphere. Example: 4749 kmNorth = N 4749. Observenorthing kilometers appearin left data display as keysare pressed.

5. INSERT/ADVANCE push-button - Press. Observe thatthe pushbutton light remainsilluminated.

6. INSERT/ADVANCE push-button -‘Press. Observe extraprecision display for presentposition northing and east-ing, to the nearest meter,appears in left and right datadisplays, respectively.

7. To load extra precision east-ing data - Press keys insequence, starting with “E”.Example: 297 m East = E297. Observe that eastingmeters appear in right datadisplay as keys are pressed.


9 . To load ext ra prec is ionnorthing data - Press keys insequence, starting with “N“

“S”. Example: 901 mNorth = N 901. Observethat northing meters appearin left data displays as keysare pressed.

3-51

TM 55-1510-221-10

NOTE

Extra precision values are always addedto normal values regardless of which key(N/S) is pressed to initiate the entry. Thenormal entry establishes the hemisphere.

10. INSERT/ADVANCE push-button - Press. Observe lati-tude and longitude data isdisplayed in UTM andINSERT/ ADVANCE push-button light extinguishes.

NOTE

The computer will convert coordinates inthe overlap area; however display valueswill reference appropriate zone.

NOTE

The “W“ key may be used to initiate east-ing entries; however computer will alwaysinterpret such entries as an “E“ input. “E“will be displayed in normal UTM display.

NOTE

Extra precision values are always addedto normal values. As an example, South4,476.995 m will display “4476S“ in nor-mal display and “995“ in extra precisiondisplay. There is no rounding between thetwo displays.

(b.) To insert geographic coordinatesof aircraft present position:

NOTE

Prior to pressing INSERT/ADVANCEpushbutton, any incorrectly loaded datacan be corrected by pressing the CLEARpushbutton and loading correct data.

1. Data selector - L/L POS.Observe that, prior to initialload, the INSERT/ADVANCE pushbutton lightis illuminated.

2. To load latitude data - Presskeys in sequence, starting

with “N” or “S“ to indicatenorth or south. Example:42°54.0' North = N 4 2 5 40. Observe that lati tudeappears in left data displayas keys are pressed.

3. INSERT/ADVANCE push-button - Press . Observepushbutton light remainsilluminated.

4. To load longitude data -Press keys in sequence, start-ing with “W” or “E” to indi-cate west or east. Example:87°54.9’ West = W 8 7 5 49. Observe that longitudeappears in right data displayas keys are pressed.

5. INSERT/ADVANCE push-button - Press. Observepushbutton light extin-guishes.

6. Data selector - DSRTK/STS.Confirm:

a. Left data display indi-cates desired trackangle in computermemory.

b. Right data display indi-cates --84, -74, -64,or --54, depending onwhich alignment statethe computer hasreached.

NOTE

After present position has been insertedand computer has advanced to state “7“,present position cannot be reinsertedwithout downmoding to STBY andrestarting alignment.

7. Data selector (CDU) -DSRTK/STS. Observe left-hand data display indicatesthe desired track in com-puter memory and right datadisplay indicates status"-194“.

3-52

NOTE

If fourth digit from right is blank, a validheading has not been stored. Proceed withnormal preflight procedure.

TM 55-1510-221-10

NOTE

8. Monitor data display for sys-tem alignment state “9“ toalignment state “8“. Observeright data display will be “---184“.

9. Monitor data display formalfunction codes. Observeif the 26V 400 Hz power isoff, “.03184“ will appear inthe right data display andWARN light illuminates. Ifmagnetic compass system isoff, “.03184“ will appear inr i g h t d a t a d i s p l a y a n dWARN light is extinguished.

10. If there are malfunctioncodes, proceed to ABNOR-MAL PROCEDURES inthis chapter.

NOTE

To achieve best accuracy, engine start andheavy loading activity should be delayeduntil entry into NAV mode.

NOTE

Waypoint data and TACAN station datamay be loaded any time after turn-on.

(3.) Verify UTM Grid Coefficients:

1. Data selector (CDU) - UTM WYPT.

2. Keys “5” and “9“ - Press simulta-neously. Observe FROM-TO dis-play is blank. Earth flatness coef-ficient appears in left display. Therelative earth radius, in meters,appears in right display.

NOTE

These values are retained from turn-on toturn-on unless changed by operator.

3. Verify that values correspond tothose required for spheroid beingused.

Values for various spheroids are listed intable 3-1.

4. If values are correct, return CDUto normal display mode bymomentarily setting data selectorto any position except UTM WYPT. If values are to be changed,continue with following steps:

(4.) Abbreviated INS Interface Test - Asrequired.

NOTE

Assuming a level aircraft, attitude indica-tors will become level during alignmentstate “8“ and remain level in all modesuntil INS is shut down. Warning indica-tors for INS attitude signals from the INSare valid while attitude sphere display islevel.

NOTE

The INS can provide test signals to theHorizontal Situation Indicator (HSI) andconnected displays. Pressing TEST switchduring STBY, ALIGN, or NAV modecauses all digits on connected digital dis-plays to indicate “8’s,” and lights theWYPT on the pilot’s HSI and the ALERTlight. Additional HSI test signals are pro-vided when INS is in ALIGN and dataselector is at any position other thanDSRTK/STS. Under those conditions,pressing TEST switch causes HSI to indi-cate heading, drift angle, and track angleerror - all at “0°" or “30°.“ At the sametime, cross track deviation is indicated at“3.75“ nautical miles (one dot) right orleft and INS-controlled HSI flags areretracted from view.

NOTE

Output test signals are supplied to theautopilot when INS steering is selected.Rotating AUTO/MAN switch to AUTOand pressing TEST during alignment fur-nishes a 15° left bank steering command.A 15° right bank steering command is fur-nished when AUTO/MAN switch is set toMAN.

3-53

TM 55-1510-221-10

NOTE

The quick test procedure may be per-formed any time after alignment State “8“is reached and prior to entry into NAV.

1. Mode selector (MSU) - ALIGN.Observe CDU displays are illumi-nated.

2. Data selector (CDU) - DSRTK/STS. Monitor right data displayuntil state “8“ (or lower) isreached. Observe right data dis-play is ---N4, where “N“ is not“9”.

3 . A U T O - M A N s w i t c h ( C D U ) -MAN.

4. Data selector - Set to any positionexcept DSRTK/STS.

5. INS - Couple to flight directorand autopilot, as applicable. Afterperforming the preceeding step,observe:

a. All lights on MSU - Checkilluminated.

b. All lights on CDU - Checkilluminated. All “8’s“ dis-played.

c. HSI - All angles 30°. Cross-track deviation bar one dotright. All INS flags retracted.

d. Flight Director/Autopilot - A15° steering command isissued.

e. Mission Control Panel - INSU P D A T E a n d N O I N SUPDATE annunciator illu-minated.

6. CDU TEST switch - Holddepressed, and rotate AUTO-MAN switch to AUTO. Observeall indications are as in step -6-except a 15° left steering com-mand is issued. On HSI, all anglesare “0°" and cross-track deviationbar is one dot left.

6 .

7 .7. CDU TEST switch - Release. If

desired, decouple INS. Observethat operation returns to normal.

(5.) To program destinations or TACANcoordinates:

NOTE

If latitude and longitude (Geographic)coordinates are being used, skip followingprocedure (5)(a.) and execute next proce-dure (5)(b.). Enter all of the data for agiven destination or TACAN before start-ing to enter data for another.

(a.) Insertion of UTM waypoint coor-dinates:

1.

2.

UTM data mayand, until finalreloaded.

3.

4.

5.

Data selector - UTM WYPT. Data displays will indi-cate last coordinates insertedinto related waypoint.

Thumbwheel - Set towaypoint n u m b e r t o b eloaded.

NOTE

be loaded in any orderentry, a value may be

To load zone and easting -Press keys in sequence, start-ing with “E”. Example: Zone16, 425 km East = E16 425.Observe that zone and east-ing in kilometers appear inthe right data display as keysare pressed.

INSERT/ADVANCE push-button - Press. Observepushbutton light is illumi-nated.

To load northing press keysin sequence, starting with“N” or “S” to indicate northor south hemisphere. Exam-ple: 4749 km North = N4749. Observe that northingkilometers appear in the leftdata display as keys arepressed.

INSERT/ADVANCE push-button - Press. Observepushbutton light remainsilluminated.

INSERT/ADVANCE push-button - Press. Observe that

3-54

TM 55-1510-221-10

TABLE 3-2. VARIOUS VALUES FOR UTM GRID COEFFICIENTS

SPHEROID

InternationalClark 1866Clark 1880EverestBesselModified EverestAustralian NationalAiryModified Airy

SOURCE: Universal Transverse Metcator

Grid Technical Manual,

TM 5-241-8, Headquarters,

Department of the Army,

30 April 1973, page 4.

Flatness Coefficient: 100 (I/f)

Relative Radius: a-6,3700.000

FLATNESS RELATIVECOEFFICIENT RADIUS

29700 8388 m29498 8206 m29346 8249 m30080 7276 m29915 7397 m30080 7304 m29825 8160 m29932 7563 m29932 7340 m

an extra precision displayrelated to resident value ofnorthing and easting, to thenearest meter, appears in leftand right data displays,respectively.

8. To load extra precision east-ing value - Press keys insequence starting with “E“.Example: 297 m East = E297. Observe that eastingmeters appear in the rightdata display as keys arepressed.

tiate the entry. The normalentry establishes the hemi-sphere.

11. INSERT/ADVANCE push-button - Press. Within 3 sec-onds computer convertsinput into latitude and longi-tude for storage in memory.The stored value is againconverted to UTM for dis-play. The INSERT/ADVANCE pushbutton lightextinguishes. Conversionroutines may cause displaysto change by up to 10 m.

NOTE


10 . To load ext ra prec is ionnorthing value - Press keysin sequence, starting with“N” or “S“. Example: 901 mNorth = N 901. Observethat northing meters appearin left data display as keysare pressed. The value isalways added to the normalvalue regardless of whichkey (N/S) is pressed to ini-

The computer will convert coordinates inoverlap area; however, data display valueswill reference appropriate zone.

NOTE

The “W“ key may be used to initiate east-ing entries; however, the computer willalways interpret such entries as an “E“input. “E“ will be displayed in normalUTM data display.

3-55

TM 55-1510-221-10

NOTE

The extra precision values are alwaysadded to normal values. As an example,South 4,476.995 m will display “4476 (S)“in the normal display and “995” in extraprecision display. In other words, there isno rounding between the two displays.

12. Repeat steps 2 through 11for each waypoint to beloaded.

NOTE

A load cycle may be terminated prior toinsertion of all four values by movingdata selector or thumbwheel.

(b.) Insertion of geographic waypointcoordinates:

1. Data selector - L/L WY PT.Data displays indicate lastcoordinates inserted into theselected waypoint.

2 . Thumbwhee l - Set towaypoint n u m b e r t o b eloaded.

3. To load latitude - Press keysin sequence, starting with“N“ or “S” to indicate northor south. Example: 42° 54.0'Nor th = N 4 2 5 4 0.Observe that INSERT/ADVANCE pushbutton lightilluminates when first key ispressed, and latitude appearsin left data display as keysare pressed.


5. To load longitude - Presskeyboard keys in sequence,starting with “W“ or “E“indicating west or eas t .Example: 87°54.9’ West =W 8 7 5 4 9. Observe thatINSERT/ADVANCE push-button light illuminateswhen first key is pressed,and longitude appears in dis-play as keys are pressed.

6. INSERT/ADVANCE push-button - Press. Observe

7.

8.

9.

10.

11.

12.

pushbutton light extin-guishes.

If desired to insert extra pre-cision coordinate data -PressINSERT/ADVANCE push-button. Observe that arc-seconds for loaded latitudeand longitude, to nearesttenth of a second, appear inleft and right data displays,respectively.

To load related arc-secondvalues for latitude - Presskeys in sequence, startingwith “N“. Example: 35.8“North = N 358.

INSERT/ADVANCE push-button - Press. Observepushbutton light extin-guishes.

To load related arc-secondvalues for longitude - Presskeys in sequence, startingwith “E“. Example: 20.1“East - E 201.


Repeat steps -2- through -11-for each waypoint to beloaded.

NOTE

In above e x a m p l e , i f INSERT/ADVANCE pushbutton was pressed, thefollowing normal display would appear:“42°54.5 (N)“ and 87°54.3(W). The extraprecision values are added to normal val-ues and normal data displays are notrounded off.

NOTE

The normal geographic coordinates mustalways be loaded prior to extra precisionvalues.

NOTE

The directions “N“ or “S” and “E“ or“W” are established during normal coor-dinate entry. Either key may be used toinitiate entry during extra precision loadsand values will be added to the extra pre-cision value without affecting direction.

3-56

NOTE

It is characteristic of the computer displayroutine to add “0.2“ arc-seconds to anydisplay of “59.9” arc-seconds. The valuein computer is as loaded by operator.

(6.) To insert TACAN coordinates:

(a.) Insertion of UTM TACAN sta-tion data:

NOTE

Prior to pressing the INSERT/ADVANCEpushbutton, any incorrectly loaded datacan be corrected by pressing CLEARpushbutton and loading correct data.

1. Data selector - UTM WYPT.

2. Keys “7“ and “9” - Presssimultaneously. Observe thatnumber of TACAN stationbeing used for navigationflashes on and off in“FROM-TO“ display anddata displays indicate coor-dinates of station selected bythumbwheel.

3. Thumbwheel - Set to num-ber of station to be loaded.Confirm thumbwheel is indetent.

4. Station “0” c a n n o t b eloaded. Observe that if num-ber of station to be loaded iss a m e a s n u m b e r o f t h eTACAN station currentlybe ing u sed , number i n“FROM-TO“ display will beset to “0“ when TACAN datais loaded.

5. To load zone and easting -Press keys in sequence, start-ing with “E“. Example: Zone16, 425 km East = E16 425.Observe that zone and east-ing in kilometers appear inthe right display as keys arepressed.

6. INSERT/ADVANCE push-button - Press. Observe thatpushbutton light is illumi-nated.

TM 55-1510-221-10

7. To load northing - Presskeys in sequence, startingwith “N“ or “S” to indicatenorth or south hemisphere.Example: 4749 km North =N 4749. Observe that north-ing kilometers appear in leftdata display as keys arepressed.


9. INSERT/ADVANCE push-button - Press. Observe thatextra precision displayrelated to the resident valueof northing and easting, tonearest meter, appears in leftand 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 normal andextra precision UTM data to be alter-nately displayed.

10.

11.

12.

To load extra precision east-ing value - Press keys insequence, starting with “E”.Example: 297 m East = E297. Observe that eastingmeters appear in right datadisplay as keys are pressed.

INSERT/ADVANCE push-button - Press. Observepushbutton remains illumi-nated.

To load ext ra prec is ionnorthing value - Press keysin sequence, starting with“N” or “S”. Example: 901 mNorth = N 901. Observethat northing meters appearin left data display as keysare pressed. The value isa lways added to -ormalvalue regardless of whichkey (N/S) is pressed to ini-tiate entry. The normal entryestablishes hemisphere.

3-57

TM 55-1510-221-10

13. INSERT/ADVANCE push-button - Press. - 10erve thatduring the next 1 to 3 sec-onds, the computer convertsinput into latitude and longi-tude for storage in memory.The stored value is againconverted back to UTM fordisplay to operator. TheINSERT/ADVANCE push-button light extinguishes.The conversion routinesmay cause data displays tochange by up to 10 m.

NOTE

The computer will convert coordinates inoverlap area; however, data display valueswill reference appropriate zone.

NOTE

The “W” key may be used to initiate east-ing entries; however, the computer willalways interpret such entries as an “E”input. “E“ will be displayed in normalUTM data display.

NOTE

The extra precision values are alwaysadded to normal values. As an example,South 4,476.995 m will display “4476 S“in normal display and “995” in extra pre-cision display. In other words, there is norounding between the two displays.

14.

15.

16.

INSERT/ADVANCE push-button - Press. Observe rightdata display indicates lastpreviously inserted altitude,and lef t da ta d isplay i sblank.

To indicate the followingload is altitude - Press keys“4” or “6“. ObserveINSERT/ADVANCE push-button light illuminates.

To load altitude in feet -Press keys in sequence.Example: 1230 ft = 1230.Observe that numbersappear in right data displayas keys are pressed.

NOTE

Altitude inputs are limited to 15,000 feet.

17.

18.

19.

20.

21.


INSERT/ADVANCE push-button - Press. Observe thatleft data display indicateslast previously insertedchannel number, and rightdisplay is blank.

To indicate following load ischannel number - Press keys"2" "8" ObserveINSERT/ADVANCE push-button light illuminates.

To load channel number -Press keys in sequence.Example: 109 = 109.Observe number appears inleft data display as keys arepressed.


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“ suf-fix.

NOTE

Degree symbol (°)should be disregardedwhen reading altitude and data display.

22. INSERT/ADVANCE push-button - Press. Observe sta-tion northing, zone, andeasting reappear.

23. Repeat steps -1- through -22-for each TACAN station.

24. To return INS to normalmode, momentarily set dataselector to UTM POS.

3-58

(b.) Insertion of geographic TACANstation data:

NOTE

Prior to pressing INSERT/ADVANCEpushbutton, any incorrectly loaded datacan be corrected by pressing CLEARpushbutton and loading correct data.

1. Data selector - L/L WY PT.

NOTE

If number of station to be loaded is sameas number of TACAN station currentlybeing used, number in “FROM-TO” dis-play will be set to “0” when TACAN datais loaded.

2. Keys “7” and “9“ - Presssimultaneously. Observe thatnumber of TACAN stationbeing used for navigationflashes on and off in“FROM-TO“ display. Datadisplays indicate coordinatesof station selected via thum-bwheel.

3. Thumbwheel - Set to num-ber of 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 northor south. Example: 42° 54.0’North = 4 2 5 4 0. Observethat INSERT/ADVANCEpushbutton light illuminateswhen first key is pressed.

5. INSERT/ADVANCE push-button - Press Observe push-button light extinguishes.

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.

7.

8.

9.

10.

11.

12.

TM 55-1510-221-10

Observe that INSERT/ADVANCE pushbutton lightilluminates when first key ispressed, and longitudeappears in data display askeys are pressed.


INSERT/ADVANCE push-button - Press. Observe thatthe arc-seconds related toloaded latitude and longi-tude, to nearest tenth of asecond, appear in left andright data display, respec-tively.

If extra precision coordinatedata is to be inserted -Presskeys in sequence, startingwith “N“, to load relatedarc-second values for lati-tude. Example: 35.8“ North= N 358.


To load related arc-secondvalues for longitude - Presskeys in sequence, startingwith “E”. Example: 20.1”East - E 201.


NOTE

In above e x a m p l e , i f INSERT/ADVANCE pushbutton were pressed, thefollowing normal display would appear:“42° 54.5 N” and “87° 54.3 W”. The extraprecision values are added to normal val-ues and normal displays are not roundedoff.

NOTE

The normal geographic coordinates mustalways be loaded prior to extra precisionvalues.

3-59

TM 55-1510-221-10

NOTE

The directions “N“ or “S” and “E” or“W“ are established during normal coor-dinate entry. Either key may be used toinitiate entry during extra precision loadsand the values will be added to extra pre-cision value without affecting direction.

NOTE

It is characteristic of the computer displayroutine to add 0.2 arc-seconds to any dis-play of 59.9 arc-seconds. The value incomputer is as loaded by operator.

13.

14.

15.

INSERT/ADVANCE push-button - Press. Observe thatright data display indicateslast previously inserted alti-tude, and left data display isblank.

To indicate the followingload is altitude - Press key“4“ or “6“. ObserveINSERT/ADVANCE push-button light illuminates.

To load altitude first - Presskeys in sequence. Example:1230 ft = 1230. Numbersappear in right data displayas keys are pressed.

NOTE

Altitude inputs are limited to 15,000 feet.

16.

17.

18.


INSERT/ADVANCE push-button - Press. Observe thatleft data display indicateslast previously insertedchannel number, and rightdata display is blank.

To indicate the followingload is channel number -Press key “2” or “8“.Observe INSERT/ADVANCE pushbutton lightilluminates.

19. To load channel number -Press keys in sequence.Example: 109 = 109. Num-bers appear in left data dis-play as keys are pressed.


NOTE

Any number will be accepted by the INS;however only stations with a channelnumber within range of 1 through 126will be used for TACAN mixing.

NOTE

The channel number has an implied “X“suffix.

NOTE

Decimal points and degree symbolsshould be disregarded when reading alti-tude and channel number displays.

21. INSERT/ADVANCE push-button - Press. Observe sta-tion latitude and longitudereappear.

22. Repeat steps -3- through -19-for each TACAN station.

23. To return INS to normal dis-play modes, momentarily setdata selector to L/L POS.

(7.) Designating fly-to destinations:

1.

2.

3.

Data selector - L/L WY PT orUTM WY PT, as required.

Waypoint thumbwheel - Selectdestination number. Observenumber of destination waypointappears in “TO” part ofFROM-TO display.

Data selector - HDG/DA.Observe present aircraft headingappears, t o n e a r e s t t e n t h o fdegree, in left data display; alsodrift angle, to nearest degree,appears in right data display.

3-60

NOTE

Navigation information is now availablefrom the INS for display on the pilot andco-pilot RMI’s and on the pilot and copi-lot HSI’s, as determined by the COURSEINDICATOR and RMI select switches.

(8.) 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 light - Monitor.Observe illumination approx 1.3minutes before reaching point forautomatic leg switch. Indicatorflashes on and off after passing awaypoint, if AUTO-MAN switchis in MAN.

(9.) Aided TACAN operation:

NOTE

If high accuracy alignmentwait for the READY NAVselecting NAV mode.

is required,light before

1. Mode selector - NAV.

2. Data selector - DSRTK/STS.

3. Key “4“ - Press. Observe rightdata display is “000004“ andINSERT/ADVANCE pushbuttonlight is illuminated.

4. INSERT/ADVANCE pushbutton- Press. Observe right data displayis "1 -XX4” and INSERT/ADVANCE pushbutton light isextinguished.

NOTE

Every 30 seconds, the INS will select nexteligible TACAN station in sequence forupdating. To be eligible, TACAN stationrange must be between 5 and 150 nm andchannel between 1 and 126.

5. Data selector - L/L WY PT orUTM WY PT.

TM 55-1510-221-10

6. Keys “7“ and “9“- Press simulta-neously. Observe channel numberof the TACAN station being usedfor navigation flashes on and off.Data displays indicate coordi-nates of station selected via thum-bwheel.

7. To monitor station selection -Observe FROM-TO data display.Observe only the number of sta-tions eligible for mixing will bedisplayed. A “0“ indicates thatnone of the 9 stations are eligiblefor selection.

8. Monitor “INS UPDATE“ annun-ciator.

NOTE

Mixing will not be annunciated if: (a)TACAN control is inappropriately set; (b)TACAN station data loaded in error; (c)aircraft look-down angle is greater than30”; (d) horizontal ground distance is lessthan two times the altitude. When 2 min-utes elapse without an update, the “NOINS UPDATE“ annunciator will illumi-nate.

9. To return INS display to normal- Set data selector to any positionexcept WYPT or DIS/TIME.

10. To monitor program of mix - Setdata selector to DSRTK/STS.(Observe Accuracy Index (AI) willdecrement to “0“.)

NOTE

To insure favorable geometry during theupdate process, the following TACAN sta-tion criteria should be observed:

11.

12.

13.

14.

One station must be at least 15nm off course.

For optimum single TACANupdating, update should continueuntil aircraft has passed the sta-tion.

F o r o p t i m u m d u a l T A C A Nupdating, use one “off-track“TACAN station and one "on-track“ station.

For optimum multi-TACAN sta-tion updating, the stations should

3-61

TM 55-1510-221-10

be evenly distributed in azimutharound the aircraft.

to nearest tenth of a degree. Release keys 3 and 9.Left display reverts to true heading.

15. Waypoint thumbwheel - Set tonumber of first TACAN station tobe used. Observe selected stationnumber is displayed on the “TO“side of FROM-TO data display.

16. Update GPS.

(e.) Ground speed: Data selector -TK/GS. Observe ground speed appears in right datadisplay to nearest knot.

(f.) Ground track angle: Data selec-tor - TK/GS. Observe ground track angle appears inleft data display to nearest tenth of a degree.

(10.) Switching from aided to unaided iner-tial operation.


2. Key “5” - Press. ObserveINSERT/ADVANCE pushbuttonlight illuminates; 000005 appearsin right data display.

(g.) Drift angle: Data selector -HDG/DA. Observe drift angle appears in right datadisplay to nearest degree.

(h.) Wind speed and direction: Dataselector - WIND. Wind direction appears in left datadisplay to nearest degree and wind speed appears inright display to nearest knot.

3. INSERT/ADVANCE pushbutton- Press. Observe INSERT/ADVANCE pushbutton lightextinguishes. Data display returnsto normal with “5“ appearing infirst digit of right display.

(i.) Desired track angle: Data selec-tor - DSRTK/STS. Observe desired track angle inright data display to nearest degree.

(j.) Track angle error: Data selector- XTK/TKE. Observe track angle error appears inright data display to nearest degree.

NOTE

Benefits of previous aiding are main-tained but no additional automaticupdates will be made.

(k.) Cross track distance: Data selec-tor - XTK/TKE. Observe cross track distanceappears in left data display to nearest nautical mile.

(11.) To obtain readouts from INS:

(l.) Distance and time to nextwaypoint: Data selector - DIS/TIME. Observe dis-tance to next waypoint, shown in “TO“ side ofFROM-TO display, appears in left data display to

NOTEnearest nautical mile. Observe time to reachwaypoint at present ground speed appears indata display to nearest tenth of a minute.

nextright

The computer is assumed to be in theNAV mode for all data displays. (m.) Extra precision geographic

(a.) System status: Data selector -DSRTK/STS. Observe numbers indicating systemstatus appear in right data display.

ent position display:

1.

pres-

(b.) Geographic present position:Data selector - L/L POS. Observe latitude and longi-tude of present position appear in left and right datadisplays, respectively. Both displays are to tenth ofa minute.

Data selector - L/L POS.Latitude and longitude ofpresent position, to nearesttenth of a minute, appears inleft and right data displays,respectively.

2.

(c.) UTM position: Data selector -UTM POS. Observe northing and zone with eastingof present position appear in left and right displays,respectively. Both displays are in kilometers.

(d.) True heading and MAG heading:Data selector - HDG/DA. Observe aircraft trueheading appears in left data display to nearest tenthof a degree. Press and hold keys 3 and 9 simulta-neously. MAG heading appears in left data display

INSERT/ADVANCE push-button - Press. Observe arc-seconds related to presentposition latitude and longi-tude, to nearest tenth of asecond, appear in left andright data displays, respec-tively.

(n.) Geographic present inertial posi-tion display.

1. Data selector - L/L WY PT.

3-62

2. HOLD pushbutton - Press.Observe HOLD pushbuttonlight illuminates, latitudeand longitude of presentinertial position to a tenth ofdegree appear in left andright data displays, respec-tively.

NOTE

While HOLD pushbutton light is extin-guished, TACAN updates are inhibited.

3. INSERT/ADVANCE push-button - Press. Observe arc-second related to presentinertial position latitude andlongitude, to nearest tenth ofa second, appears in left andright data displays, respec-tively.

4. HOLD pushbutton - Press.Observe INS returns to nor-mal operation and HOLDpushbutton light extin-guishes.

display:(o.) UTM present inertial position

1. Data selector - UTM WYPT.

2. HOLD pushbutton - Press.Observe HOLD pushbuttonlight illuminates. Northingand zone with easting of thepresent inertial position inkilometers appear in left andright data displays, respec-tively.

NOTE

While HOLD pushbutton light is illumi-nated, TACAN updates are inhibited,

3. INSERT/ADVANCE push-button - Press. Observe extraprecision values related topresent inertial positionnorthing and easting, tonearest meter, appear in leftand right data displays,respectively.

4. HOLD pushbutton - Press.Observe INS returns to nor-

TM 55-1510-221-10

ma1 operation and HOLDpushbutton light extin-guishes.

(p.) Distance and time to waypointother than next waypoint:

1. WYPT CHG pushbutton -Press. Observe WYPT CHGand INSERT/ADVANCEpushbutton light illuminates.

2. Key “0“ - Press. Observe“FROM“ side of FROM-TOdata display changes to “0“.

3. Key corresponding todesired waypoint - Press.O b s e r v e “ T O “ s i d e o fFROM-TO da ta d i sp laychanges to desired waypointnumber.

NOTE

Do not press INSERT/ADVANCE push-button. This would cause an immediateflight plan change.

4.

5.

Data selector - DIS/TIME.Observe distance to desiredwaypoint appears in left datadisplay to nearest nauticalmile. Time to reach desiredwaypoint at present ground-speed appears in right datadisplay to nearest tenth of aminute.

CLEAR pushbutton - Press.Observe INS returns to nor-mal operation. ObserveINSERT/ADVANCE andWYPT CHG pushbuttonlights extinguish. Waypointsdefining current navigationleg appear in FROM-TO dis-play.

(q.) Distance and time between anytwo waypoints:

1. WYPT CHG pushbutton -Press. Observe WYPT CHGand INSERT/ADVANCEpushbutton lights illuminate.

2. Keys corresponding todesired waypoints - Press insequence. Observe desiredwaypoint numbers appear in

3-63

TM 55-1510-221-10

FROM-TO data display askeys are pressed.

NOTE

Do not press INSERT/ADVANCE push-button. This would cause an immediateflight plan change.

3. Data selector - DIS/TIME.Observe distance betweendesired waypoints appears inleft data display to nearestnautical mile. Time to travelbetween desired waypointsat present ground speedappears in right data displayto nearest tenth of a minute.

4. CLEAR pushbutton - Press.Observe INS returns to nor-mal operation. ObserveWYPT CHG and INSERT/ADVANCE pushbutton lightextinguishes. Waypointsdefining current navigationleg appear in FROM-TOdata display.

(r.) Distance to any TACAN station:

1. Data selector - DIS/TIME.Observe distance to nextwaypoint to nearest nauticalmile is in left data display.Time to next waypoint tonearest tenth of a minute isin right data display.

2. Keys “7“ and “9“ - Presssimultaneously. Observenumber of TACAN stationbeing used for navigationflashes on and off inFROM-TO display. Distanceto TACAN station to nearestnautical mile is in left datadisplay. T i m e t o nextwaypoint is in right data dis-play.

3. If in aided TACAN opera-tion - Monitor display.Observe station number isselected every 30 seconds.

4. If not in aided TACANoperation - Perform steps 5through 7.

5. WYPT CHG pushbutton -Press. Observe INSERT/

ADVANCE and WYPTCHG pushbutton lights illu-minate. Station numberflashing discontinues.

6. Key indicating desiredTACAN station number -Press. Observe number willappear in left digit locationof FROM-TO data display.

NOTE

If wrong key is pressed, press CLEAR;displays will revert to that indicated instep 2.

7. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE andWYPT CHG pushbuttonlights extinguish. The loadeddigit will appear in rightposition of FROM-TO dis-play and will be flashing onand off. Distance to that sta-tion to nearest nautical mileappears in left data display.The right display continuesto d isplay t ime to nextwaypoint.

8. Data selector - WIND,momentarily. Returns INSto normal display mode.

NOTE

If in aided TACAN operation and if thedesired station is not being selected, exitaided operation per procedure: “Switch-ing From Aided to Unaided InertialOperation“, perform steps 1 thru 8, andthen return to aided operation per proce-dure: “Aided TACAN Operation“.

(s.) Coordinates of any waypoint:

1. Data selector - L/L WY PTor UTM WY PT.

2. Waypoint thumbwheel - Setdesired waypoint. Observefollowing:

a. L/L WY PT: latitudeand l o n g i t u d e o fdesired waypoint, to at e n t h o f a m i n u t e ,appear in left and right

3-64

TM 55-1510-221-10

data displays respec-tively.

b. UTM WY PT: North-ing and zone with east-ing of desiredwaypoint, to a kilome-ter, appear in left andr ight da ta d isp laysrespectively.

3. INSERT/ADVANCE push-button - Press. Observe thefollowing:

a. L/L WY PT: The arc-seconds r e l a t e d t odesired waypoint lati-tude and longitude, toa tenth of an arc-second appear, in leftand right displaysrespectively.

b. UTM WY PT: Theextra precision displayr e l a t e d t o desiredwaypoint northing andeas t ing , in meters,appear in left and rightdata displays respec-tively.

NOTE

L/L WY PT: A coordinate is the additionof values for degrees, whole minutes andseconds.

Example: W 87° 54’ 58.6“ = 87°54.9Wand 58.6.

UTM WY PT: A coordinate is the addi-tion of the values for kilometers andmeters.

Example: S 2,474,706m = 2474S and706.

(t.) TACAN station data:

1. Keys “7” and “9” - Presssimultaneously.

2. Waypoint thumbwheel - Setto desires TACAN station.Observe number of TACANstation being used for navi-gation flashes on and off.

a. L/L WY PT: Latitudeand l o n g i t u d e o fdesired TACAN sta-

tion, to tenth of min-ute, appears in left andright data displays,respectively.

b. UTM WY PT: North-ing and zone with east-ing of desired TACANstation, to a kilometer,appear in left and rightdata displays, respec-tively.

3. INSERT/ADVANCEPUSHBUTTON - Press.Observe the following:

a. L/L WY PT: The arc-seconds r e l a t e d t odesired TACAN sta-tion, to tenth of an arc-second, appear in leftand right data displaysrespectively.

b. UTM WY PT: Theextra precision displayr e l a t e d t o desiredTACAN station north-ing and easting, inmeters, appear in leftand right data displaysrespectively.

NOTE

Direction is indicated in normal data dis-plays.

4. L/L WY PT: A coordinate isthe addition of values fordegrees, whole minutes, andseconds.

5. Example: W 87° 54’ 58.6“will be displayed as “87°54.9W” and 58.6”.

6. UTM WY PT: A coordinateis the addition of values forkilometers and meters.

7. Example: S 2,474,706 m willbe displayed as “2474 S”and “706“.

8. INSERT/ADVANCE push-button - Press. ObserveTACAN station altitude, infeet, will appear in right datadisplay; degree symbol anddecimal points should be

3-65

TM 55-1510-221-10

disregarded. Left data dis-play is blank.

9. INSERT/ADVANCE push-button - Press. ObserveTACAN sta t ion channelnumber, in whole numbers,will appear in left data dis-play; degree symbol and dec-imal point should be disre-garded. Right data display isblank.

NOTE

If INSERT/ADVANCE pushbutton ispressed, the normal coordinates indicatedin step -3- will be displayed.

NOTE

Waypoint thumbwheel may be moved atany time and normal coordinates for newTACAN station will be displayed.

10. Data selector - Momentarilyto any position other thanL/L WY PT, UTM WY PTor DIS/TIME. (Returns INSto normal operation.)

(u.) Magnetic heading.

1. Data selector - HDG/DA.Observe true heading tonearest tenth degree appearsin right data display.

2. Keys “3“ and “9” - Presssimultaneously and hold.Observe magnetic heading tonearest tenth of a degreeappears in left data display.Drift angle continues to bedisplayed in right data dis-play.

3. Keys “3“ and “9” - Release.Observe left data displayreverts to true heading.

(12.) INS updating:

(a.) Normal geographic present posi-tion check and update:

1. Data selector - L/L POS.Observe latitude and longi-

2.

tude of present positionappear in left and right datadisplays, respectively.

Illuminated HOLD pushbut-ton - Press. Observe latitudeand longitude in data dis-plays freeze at values presentwhen HOLD pushbutton ispressed.

NOTE

While HOLD pushbutton light is illumi-nated, TACAN, GPS and data linkupdates are inhibited.

3. Keys - Press in sequence toload latitude of position ref-erence, starting with “N“ or“S” to indicate north orsouth. Example: 42°54.0'north = N 4 2 5 4 0.Observe INSERT/ADVANCE pushbutton lightilluminates when first key ispressed, and latitude appearsin left data display as keysare pressed.

4. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE push-button light remains illumi-nated, and previous value oflatitude reappears.

5. Keys - Press in sequence toload longitude of positionreference, starting with “W”or “E“ to indicate west oreast. Example: 87°54.9’ west= W 8 7 5 4 9. Observe lon-gitude appears in right datadisplay as keys are pressed.

6. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE andHOLD pushbutton lightsremain illuminated. Northposition error and east posi-tion error, in tenth of a nau-tical mile, will appear in leftand right data displays,respectively.

3-66

TM 55-1510-221-10

NOTE

If WARN light illuminates, proceed tostep 7; otherwise proceed to step 8.

7. Data selector - DSRTK/STS.Observe action code “02”and malfunction code “49“.This indicates that the radialerror between the loadedposition and the INS posi-tion exceeds 33 nauticalmiles. Operator must evalu-ate possibility that eitherINS is in error or referencepoint position is incorrect. Itis possible to force INS toaccept updated position bysetting data selector to L/LPOS and proceeding to step8).

8. If displayed values arewithin tolerance, pressHOLD pushbutton to returnINS to normal operation. Ifone or both values are out oftolerance, proceed to step 9.

9. Key “2” - Press. Observe leftdata display is “00000 N”;INSERT/ADVANCE andHOLD pushbutton lights areilluminated.

10. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE andHOLD pushbutton lightsextinguish. Present positionappears in data displays. Pres-e n t p o s i t i o n c h e c k a n dupdate is complete.

N O T E

Within 30 seconds, computer will processcorrection and revised present positionwill appear in data display. If AI prior toposition update is 1 or greater, computerwill accept over 95 percent of correctionshown in difference display. If AI is “0“,amount of correction accepted will be lessand is a function of time in NAV modeand number of updates which have beenmade.

(b.) Extra precision geographic pres-ent position check and update:


2. Key “2” - Press. ObserveINSERT/ADVANCE push-button light illuminates,“000002” appears in rightdata display.

3. INSERT/ADVANCE push-button - Press. Observe rightdata display is “1-XX2“,INSERT/ADVANCE push-button light is extinguished,and any TACAN, GPS ordata link updating is discon-tinued.

4. Data selector - L/L POS.Observe latitude and longi-tude of present positionappears in left and right datadisplays, respectively.

5. HOLD pushbutton - Press(when aircraft passes overknown position reference.)Observe HOLD pushbuttonlight illuminates. Latitudeand longitude in data dis-plays freeze at values presentwhen HOLD pushbuttonwas pressed.

6. Load latitude by pressingkeys in sequence, startingwith “N“ or “S” to indicatenorth or south. Example:42°54.0’ North = N 4 2 5 40. Observe latitude appearsin left data display as keysare pressed.

7. INSERT/ADVANCE push-b u t t o n - Press ObserveINSERT/ADVANCE andHOLD pushbuttons remainilluminated.

8. Load longitude by pressingkeys in sequence, startingwith “W” on “E” indicatingw e s t o r east. Example:87°54.9’ West = W 8 7 5 49.

9. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE andHOLD pushbutton lightsremain illuminated.

3-67

TM 55-1510-221-10

10.

11.

12.

13.

INSERT/ADVANCE push-button - Press. Observe arc-seconds related to presentposition latitude and longi-tude, to nearest tenth of asecond, appear in left andright data displays, respec-tively.

Load related arc-second val-ues for latitude in sequence,starting with “N“. Example:35.8° North = N 358.

INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE andHOLD pushbuttons remainilluminated.)

Load related arc-second val-ues for long i tude i nsequence, starting with “E“.Example: 20.1° East = E201.

NOTE

Extra precision values are added to nor-mal values and normal displays are notrounded off.

NOTE

Normal latitude-longitude coordinatesmust always be loaded prior to extra pre-cision values.

NOTE

Directions “N” or “S“ and “E” or “W”are established during normal coordinateentry. Either key may be used to initiateentry during extra precision loads andvalues will be added to extra precisionvalues without affecting direction.

NOTE

It is characteristic of the computer displayroutine to add 0.2 arc-seconds to any dis-play of 59.9 arc-seconds. Value in com-puter is loaded by operator.

14. Proceed to step 6 in proce-dure: “Normal GeographicPresent Position Check andUpdate.“

update:(c.) UTM present position check and

UTM data mayand, until finalreloaded.

1.

2.

NOTE

be loaded in any orderentry, a value may be

Data selector - UTM POS.Observe UTM coordinatesof present position appear indata displays.

HOLD pushbutton - Press(when aircraft passes overknown position reference.)Observe HOLD pushbuttonlight illuminates. Coordi-nates in data display freezeat values present whenHOLD pushbutton waspressed.

NOTE

While HOLD pushbutton light is illumi-nated, TACAN, GPS and data linkupdates are inhibited.

3.

4.

5.

6.

Load zone and easting bypressing keys in sequence,starting with “E”. Example:Zone 16, 425 km East =E16 425. Observe zone andeasting in kilometers appearin right data display as keysare pressed.

INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE push-button light remains illumi-nated.

Load northing by pressingkeys in sequence, startingwith “N” or “S” to indicatenorth or south hemisphere.Example: North 4749 km =N 4749. Observe northingkilometers appear in leftdata display as keys arepressed.

INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE push-button light remains illumi-nated.

3-68

TM 55-1510-221-10

7. INSERT/ADVANCE push-button - Press. Observe extraprecision display related topresent position northingand easting, to nearestmeter, appears in left andright data displays, respec-tively.

8. Load extra precision eastingvalue by pressing keys insequence, starting with “E“.Example: 297 m East = E297. Observe easting metersappear in right data displayas keys are pressed.

9. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE push-button light remains illumi-nated.

10. Load extra precision north-ing value by pressing keys insequence, starting with “N“or “S”. Example: 901 mNorth = N 901. ObserveNorthing meters appear inleft data display as keys arepressed. The value is alwaysa d d e d t o n o r m a l valueregardless of which key(N/S) is pressed to initiateentry. Normal entry estab-lishes the hemisphere.

NOTE

The “W” key may be used to initiate east-ing entries; however, the computer willalways interpret such entries as an “E“input.

NOTE

The extra precision values are alwaysadded to normal values.

NOTE

Any data inserted when HOLD pushbut-ton light is not illuminated will berejected by computer.

11. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE andHOLD pushbutton lightsremain illuminated. North

position error and east posi-tion error in kilometers willappear in left and right datadisplays, respectively.

12. If WARN light illuminates,proceed to step 13; other-wise proceed to step 9 inprocedure: “Extra PrecisionGeographic Present PositionCheck and Update.“

13. Data selector - DSRTK/STS.Observe action code “02“and malfunction code “49“.This indicates radial errorbetween loaded position andINS position exceeds 62kilometers. Operator mustevaluate possibility that INSis in error or reference pointposition is incorrect. It ispossible to force INS toaccept updated position bysetting data selector to UTMPOS and proceeding to step10 of procedure: “Extra Pre-cision Geographic PresentPosition Check

14. If updating is to be rejected- Press HOLD pushbutton.Observe HOLD andINSERT/ADVANCE push-button lights extinguish. INSreturns to normal operation.

(d.) Position update eradication:

NOTE

This procedure is not considered com-mon. Its use is limited to those timeswhere an operational error has resulted inan erroneous position fix.


2. Key “1“ - Press. ObserveINSERT/ADVANCE push-button light illuminates,000001 appears in right datadisplay.

3. INSERT/ADVANCE push-button - Press. ObserveINSERT/ADVANCE push-button light extinguishes.Within 30 seconds, data dis-plays return to normal with“0“ (normal inertial mode)

3-69

TM 55-1510-221-10

in last digit of right display.AI will be set to approxi-mately three times the num-ber of hours in NAV.

(13.) Flight course changes.

(a.)tion:

Manual flight plan change inser-

1. WYPT CHG pushbutton -Press. Observe WYPT CHGand INSERT/ADVANCEpushbuttons illuminate.

2. Select new FROM and TOwaypoints bysponding keys

pressing corre-

3. WYPT CHG pushbutton -Press. Observe newwaypoint numbers appear inFROM-TO data displays askeys are pressed.

NOTE

Selecting zero as FROM waypoint willcause desired track to be defined by com-puted present position (inertial presentposition plus fixes) and TO waypoint.

4. INSERT/ADVANCE push-button - Press. ObserveWYPT CHG and INSERT/ADVANCE pushbuttonsextinguish.

NOTE

Waypoint zero always contains rampcoordinates if no manual flight planchanges are made. When a manual flightplan change is made, present position atinstant of insertion is stored in waypointzero.

(14.) After landing procedures:

If INS will be unattended for an extendedperiod, it should be shut down.

Do not leave INS operating unless aircraftor ground power and cooling air are avail-able to system.

NOTE

The INS may be shut down, downmodedto STBY or ALIGN mode, or operated inthe navigation mode after landing. Thedetermining factor in choosing course ofaction is expected length of time beforethe next takeoff.

NOTE

Do not tow or taxi aircraft during INSalignment. Movement during alignmentrequires restarting alignment.

(a.) Transient stops.

NOTE

Action to be taken during a transient stopdepends upon time available and onavailability of accurate parking coordi-nates (latitude and longitude.)

1. Realignment - INS operat-ing.

(Recommended if sufficient time and accurateparking coordinates are available.)

NOTE

INS can be downmoded to perform arealignment and azimuth gyro calibration.Alignment to produce an alignment statenumber of “5” can be accomplished inapproximately 17 minutes. During the 17minute period, an automatic azimuthgyro recalibration is determined on basisof difference between inertial presentposition before downmoding and insertedpresent position. To obtain further refine-ment of azimuth gyro drift rate, calcu-lated on basis of newly computed errordata, INS can be left in alignment modefor a longer period, allowing the align-ment state number to attain some valuelower than “5“.

2. Data selector - STBY, thento ALIGN.

3. Present position coordinates- Insert, according to proce-dure: “Geographic PresentPosition I n s e r t i o n ” o r“UTM Present Position

3-70

TM 55-1510-221-10

4. Realignment - INS shut-down. Perform completealignment procedures.

5. Position update. Recom-mended if time is not avail-able for realignment.

NOTE

Perform position update using parkingcoordinates in accordance with proce-dure: “Insertion of Geographic WaypointCoordinates.“ If parking coordinates arenot available, proceed as follows:

Continue operation in NAV, if INS accu-racy appears acceptable.

Perform position update using best esti-mate of parking coordinates.

6. Downmoding to standby:Mode selector - STBY.

NOTE

INS can be downmoded to standby opera-tion which will maintain navigation unitat operating temperature with gyro wheelsrunning. INS is downmoded to standby asfollows:

Do not leave INS operating unless aircraftor ground power and cooling air are avail-able to system.

7. Shutdown: Mode selector -OFF.

N O T E

INS will retain inertial present positiondata computed at time INS is down-moded. This value is compared with pres-ent position inserted for next alignmentand difference is used to determine azi-muth gyro drift rate.

e. Abnormal Procedures:

(1.) General. INS contains self-testing fea-tures which provide one or more warning indica-tions when a failure occurs. The WARN light on thCDU provides a master warning for most malfunc-tions occuring in the navigation unit. Malfunctionsin the MSU or CDU will normally be obviousbecause of abnormal indications of displays andlights. A battery unit malfunction will shut downINS when battery power is used.

(2.) Automatic INS shutdown.

(a.) Overtemperature. An overtemp-erature in navigation unit will cause INS to shutdown (indicated by blank CDU displays) whenmode selector is at STBY or ALIGN during groundoperation. The WARN light on CDU will illuminateand will not extinguish until mode selector is rotatedto OFF. The cooling system should be checked andcorrected if faulty. If cooling system is satisfactory,navigation unit should be replaced.

(b.) Low battery charge. A low bat-tery unit charge will cause INS to shut down whenINS is operating on battery unit power. Both WARNlight on CDU and BAT light on MSU will illumi-nate and not extinguish until the mode selector is setto OFF. The battery unit should be replaced whenthis failure occurs.

(c.) Interpretation of failure indica-tions. It is important to be able to correctly interpretfailure indications in order to take effective action.Failure indications are listed below under two maincategories: WARN light illuminated, and WARNlight extinguished. Under each of these categoriesare listed other indications which will give the oper-ator sufficient information to take action.

1. WARN light illuminated.Take the following action:

a. If action codes 01, 02,03, 04, 05 are displayed- See table 3-2.

b. No action or malfunc-tion codes displayed -Indicated computerfailure.

c. Improper displays -Indicates NU computerfailure.

2. WARN light extinguished. IfCDU displays are blank,incorrect or frozen - CDUfailure is indicated.

3-71

TM 55-1510-221-10

NOTE

It is not possible to load displays from thekeyboard. A temporary failure of anumerical key may prevent data loading.If a number cannot be loaded into lati-tude or longitude displays, after pressing/wiggling the key several times, the causemay be the momentary hang-up ofanother key. To identify the faulty key,rotate the data selector to DSRTK/STS.The right digit on right display will indi-cate suspect key. Press and release suspectkey several times. To test whether thekeyboard problem is corrected, try press-ing any other numerical key. Its numbershould now appear as the right digit. Ifthis test is successful, press the CLEARkey and return data selector to originaldata loading position. Otherwise, a CDUfailure is indicated.

minated.(d.) CDU BAT indicator light is illu-

Operation on battery is an indication thatthere may may be no aircraft power toblower motor with resultant loss of cool-ing. The INS can operate only a limitedtime (normally 15 minutes) on batterypower before a low voltage shutdown willoccur. Then, immediate corrective actionmust be taken.

NOTE

CDU BAT indicator will illuminate for12 seconds in alignment State “8“ (about5 minutes after turn-on). This is normaland indicates a battery test is in progress.No corrective action is required.

NOTE

During ground operation, it is recom-mended that operation on battery powernot exceed 5 minutes.

1. To determine correctiveaction: (Monitor CDU dis-plays while rotating theCDU selector switch.)

a. If displays are frozen(do not change while

data selector is beingrota ted) problem isnormally in the naviga-tion unit.

b. If displays respond nor-mally to the data selec-tor, the problem is nor-mally absence of 115VAC power to INS.

(e.) For corrective action: Check toassure proper settings of following switches and cir-cuit breakers essential to INS operation:

1. Overhead circuit breakerpanel (fig. 2-26) - Circuitbreakers in:

a. AVIONICS MASTERCONTR

b. INS CONTROL

c. AVIONICS MASTERPWR No.1

d. AVIONICS MASTERPWR No.2

2. Overhead control panel (fig.2-12): INVERTER No. 1 orINVERTER No.2 switch -ON (either).

3. Mission control panel (fig.4-1):

a. #1 INV or #2 INVswitch - ON.

b. Bus cross tie switch -ON/AUTO.

4. Mission AC/DC Power Cab-inet (fig. 2-2): INS AC PWRcircuit breaker - In.

NOTE

CDU BAT indicator should extinguishafter above corrective action. If it remainsilluminated, INS will eventually shutdown when battery voltage drops belowapproximately 19VDC. Flight crewshould prepare for shutdown.

(3.) Malfunction indications and proce-dures: Table 3-3 details the procedure for a Malfunc-tion Code Check. Table 3-4 lists a number of mal-function indications which occur under given modesof operation. Follow procedure given. Table 3-5details action codes and recommended action. Table

3-72

TM 55-1510-221-10

3-6 lists failed test symptoms by malfunction codesand lists codes for recommended actions.

Table 3-3. Malfunction Code Check

stepIndication Control Operation

12 Data Selector Rotate to DSRTK/STS

3 TEST switch Press and release

Indicator orDisplay

WARN light LightsRH data display Action code second and third

digitsRH data display Lowest number malfunction code

which has occurred since this

Indication

procedure was performed replacesaction code.

4

5

Repeat step 3 repeatedly, recording all malfunction codes until second and third digits again indicatean action code or go blank. Refer to Table 3-4 for action codes and recommended action and to Table3-5 for malfunction code definition.If WARN light extinguishes and two digits go blank, failure was intermittent and has been cleared. If dig-its do not go blank, perform action according to displayed recommended action code.

3-73

TM 55-1510-221-10

Table 3-4. Malfunction Indications and Procedures

Mode of MalfunctionOperation Indication Procedure Probable Cause

STBY or ALIGN WARN on, CDU blank(DIM control clockwise),MSU BAT off

1. Rotate MSU modeselector OFF.

2. Check aircraftcooling system andcorrect if faulty.

3. Realign INS.

Automaticshutdowncaused byover-temperature.

STBY, ALIGN, NAV WARN on, MSU BAT on, 1. Rotate MSU mode Loss of INSCDU blank. selector OFF. power and low

battery2. Insure all switches Unit (BU).

and circuit breakersapplicable to INSoperation are setproperly.

3. If in flight, rotateMSU mode selectorOFF.

4. If on ground, replacebattery unit. Batteryunit test may be by-passed by rotatingmode selector toOFF, then to NAVand reloadingposition coordinates.When INS advancesto alignment State 7(PI=7) rotate modeselector to ALIGN.

STBY, ALIGN, or NAV WARN on, CDU is Perform Malfunction Navigationoperating Code Check as failure or

described in Table 3-2. interfacingsystem problem.

3-74

TM 55-1510-221-10

Table 3-5. Action Codes and Recommended Actions

Code Recommended Action

0102

03

0405

Shut down INS.Watch for degradation (NAV), During ground operation, downmode to STBY and restartalignment.INS may be used for navigation. One or more analog outputs are not functioning properly. Check26 VAC circuit breakers, HSI and autopilot.Downmode to STBY and restart alignment (ground operation only).Correct problem in interfacing system (could be INS). Will not seriously affect performance.

3-75

TM 55-1510-221-10

Table 3-6. Recommended Codes

MalfCode Failed Test

10 Invalid heading11 GR/CS program pin connected in error12* Canned altitude profile in use (input altitude invalid)13 Y velocity change14 X velocity change15 Torque limited16 Invalid pitch and roll17 Invalid magnetic heading16 Excessive saturation time

20* Bearing to waypoint22* Bearing to waypoint23* Drift angle24* Steering converter25* True heading converter26* XTK converter27* Tick mark to fast31 Ground speed32 Memory parity33 Azimuth stabilization loop34 Inner roll stabilization loop35 Pitch stabilization loop36 Accelerometer loop37 Z platform overtemperature38 XY platform overtemperature40 Heading error42 Drift angle !gt!45°44 Azimuth gyro drift ratye45 Gyro scale factor or loaded altitude47 15-second loop49 Fix measurement too large51* Excessive wind54* Incomplete conversion from UTM to L/L57 XY platform rotation rate59 600 millisecond loop60 X or Y sample and hold change62 XY platform rotation rate63 CDU self-checks

*Failed test does not illuminate WARN light on CDU.

Modes of RecommendedOperation Action Code

ALIGN 04ALIGN 01ALIGN, NAV 05NAV 02NAV 02ALIGN, NAV 02ALIGN, NAV 05ALIGN, NAV 05ALIGN 04ALIGN, NAV 03ALIGN, NAV 03ALIGN, NAV 03ALIGN, NAV 03ALIGN, NAV 03ALIGN, NAV 03STBY 01NAV 02STBY, ALIGN, NAV 02ALIGN, NAV 01ALIGN, NAV 01ALIGN, NAV 01ALIGN, NAV 01NAV 01NAV 01ALIGN 04NAV 02ALIGN 02ALIGN 04NAV 02NAV 02ALIGN, NAV 05STBY, ALIGN, NAV 05ALIGN 02STBY, ALIGN, NAV 02ALIGN 04NAV 02STBY, ALIGN, NAV 02

3-76

TM 55-1510-221-10

3-30A. GLOBAL POSITIONING SYSTEM(AN/ASN-149, (B)3).

a . Description. Complete provisions are in-stalled for a global positioning system (GPS). TheGPS is used to provide updated position informa-tion to the inertial navigation system. The GPS sys-tem consists of a control/display unit, receiver, ob-server headset and GPS key panel, antennaelectronics unit, and an antenna.

ed by the data select switch. If more than one pageis available a double arrow is displayed in the lowerright comer of the display. Pressing the slew key willaccess the next page. Repeated pressing of the slewkey will return the display to the first page after thelast page has been accessed.

(1.) Control/display unit (CDU). The con-trol/display unit (fig. 3-25A), located on the elec-tronics rack in the cabin, accomplishes all displayand control functions necessary for the operation ofthe GPS receiver.

(i.) Data selector switch. For all dataselector switch positions there are two modes of dis-played data:

1. Destination mode (active waypoint asdestination)

2. Waypoint (WP) examine mode (anywaypoint)

(2.) Observer Headset and GPS Key Panel.The observer headset and GPS key panel (fig. 3-25B), located on the electronics rack in the cabin,contains a headset connector and GPS key and load-ing controls.

Pressing the WP key switches the CDU betweenthe two modes.

The 10 position data select switch is used to se-lect the type of information to be displayed on the

b. GPS Controls, Indicators, and Functions.

(1.) GPS control/display unit (fig. 3-25A).

(a.) Line selection keys. Four line se-lection keys, located to the left of the CDU displayscreen, are used to initiate and terminate data en-tries, and to select various system options.

CDU:

POS.

MSN.

OPT.

(b.) Display screen. System informa-tion is shown on the cathode ray tube display screen.The display screen can show four lines of text with13 alphanumeric characters on each.

STAT.

VAR-DTM.

(c.) Mode selector switch. The four-position mode selector switch, placarded PULLOFF, INIT, NAV, and PULL TEST, is used to selectthe operating mode of the GPS system.

ERR.

WIND.

DIS-TG.

Position data is displayed.

Mission data is displayed.

Option data is displayed. Six pagesof information pertaining to theGPS receiver are made availablewhen the OPT position is selected.

Status data is displayed.

Magnetic variation and map datumdata is displayed.

Error data is displayed.

Wind data is displayed.

Distance and time to go data is dis-p l ayed .

(d.) Display brightness control. Acontrol knob placarded BRT is provided to controlthe brightness of the cathode ray tube display screen.Clockwise rotation of the control increases bright-ness.

TRK-GS. Track and ground speed data is dis-played.

DTK-VA. Desired track and vertical angledata is displayed.

(e.) Data entry keys (0 through 9). (j.) Waypoint key. The waypointThe data entry keys are used to enter alphanumeric key, placarded WP, is used to enter and examinedata. waypoint data.

(f.) USE LTR key. The use letterkey, placarded USE LTR, is used to select alphabet-ic prompt in free format data entry. The USE LTRkey terminates alphabetic entry when pressed.

(k.) Mark key. The MARK key isused for MARK and FREEZE functions.

(2.) Observer Headset and GPS Key PanelControls, Indicators, and Functions fig. 3-25B).

(g.) Clear key. The clear key, plac-arded CLR, is used to clear erroneous data entryand message displays.

(h.) Slew key. The slew key is used toaccess additional pages within a data display select-

(a.) Interphone hot microphone, nor-mal, key radio switch. This switch, placarded INT-PH HOT MIC-NORM-KEY RADIO allows selec-tion of hot microphone intercom, normal, and keyradio positions.

Change 2 3-76A

TM 55-1510-221-10

Figure 3-25A. GPS Control/Display Unit (CDU)

3-76B Change 2

TM 55-1510-221-10

Figure 3-25B. Observer Headset and GPS Key Panel

Change 2 3-76C

TM 55-1510-221-10

(b.) Observer headset connector. Al-ows connection of observer headset.

(c.) GPS Zeroize switch. Actuatinghe guarded switch, placarded ZEROIZE GPS, willdeclassify the GPS receiver.

GPS key.(d.) GPS key connector. Connects

(e.) Load switch. This pushbuttonswitch initiates loading process.

(f.) Load status indicator light. Illu-ninates to indicate load status.

(g.) Dust cap. Covers GPS key con-lector when not in use.

(h.) Observer microphone connector.Connects observer microphone.

(i.) GPS time of day switch. Thispushbutton switch is used to transmit GPS time ofday to Have Quick II radios.

c. GPS System Modes of Operation.

(1.) Off Mode. When the PULL OFFnode has been selected, power is removed from thesystem, except panel lighting.

NOTE

Critical memory and other circuits whichcannot be turned off remain powered bybatteries in the receiver.

(2.) Initialize Mode. When the INIT (ini-tialize) mode has been selected, position and time,estimates can be entered via the keypad. Waypointdata may be entered and examined, and option se-lections made. No navigation functions can be per-formed.

(3.) Navigation Mode. Selection of theNAV (navigation) mode enables normal GPS func-tions (satellite tracking and navigation), includingdata transfer to and from other aircraft systems.

(4.) Test Mode. Selection of the PULLTEST mode initiates a full command test of GPSuser equipment for line replaceable unit (LRU) faultidentification and isolation.

d. GPS Normal Operation.

(1.) GPS Start Procedures. The GPS mustbe initialized prior to being used for navigation.There are three types of start: normal, quick, andcold. A position estimate, time estimate, and alma-nac (or ephemeris) data are required for normal

3-76D Change 2

start. A quick start uses stored position, time, andrecent ephemeris information. A cold start is usedonly when the GPS is unable to perform a normalstartup.

(a.) GPS Normal Start.

1. Mode selector switch - INIT.When built-in-test is com-plete the display will showdata corresponding to thedata selector switch position.

NOTE

Data display will not be illuminated forabout 30 seconds after GPS has beenturned on. Ensure that the display bright-ness control has been set to the full clock-wise position to receive the INIT display,then adjust as desired.

NOTE

If the GPS has been OFF for more than30 seconds when INIT mode was selected,the set will perform the initial built-in-testwhich takes approximately 30 seconds.

2.

3.

4.

5.

Data selector switch - POS.If ENTER POS message isdisplayed press line selectkey 3 next to message. Posi-tion must be entered.

Displayed position - Check.Verify or enter new updatedposition and altitude as re-quired.

Data selector switch - TRK-GS. Verify correct track andgroundspeed are displayed.If not valid, enter correctvalues.

Slew key-Press. Enter cur-rent time, year, and day ofyear on page 2.

NOTE

Prior to next step, ensure all required ini-tialization data has been entered correct-ly, as they cannot be changed after selec-tion of NAV mode.

6. Mode selector switch - NAV.GPS will begin to search folsatellite signals.

NOTE

If COLD alternates with the figure ofmerit display, the GPS is performing acold start.

7. Data selector switch - STAT.

NOTE

The number of satellites (SAT) being ac-quired and tracked can be observed. Esti-mated position error (EPE) and figure ofmerit (FM) can be monitored. The GPSwill be ready for use when SAT 3 or SAT4 is displayed on STAT page 1.

8. Select page 2 of STAT.Check almanac age (ALM).If greater than 5000 hours,force a cold start.

9. While the GPS is acquiringsatellites, periodically checkSTAT page 1 for SAT 3 orSAT 4 message. Figure ofmerit (FM) is another indi-cation of a converging posi-tion fix and can be directlymonitored from page 1 ofany data selection, whereFM alternates with the sys-tem map datum and otheralerts.

10. SAT 3 or SAT 4 should bedisplayed within five min-utes. If not, check that posi-tion, time, track, andgroundspeed have been en-tered correctly. Also checkthat satellites are available.If all information is correctand satellites are available,force a cold start.

(b.) GPS Quick Start.

1. Mode selector switch - Set toNAV directly from OFF. Af-ter power-on test has beencompleted, the GPS uses ve-locity estimates from the air-craft’s sensors (if available).If velocity is not availablefrom the aircraft, zero veloc-ity is assumed. If positionand time are not availablefrom the aircraft, the posi-tion estimate from GPS

2.

3.

4.

TM 55-1510-221-10

memory is used, and the in-ternal low power time source(LPTS) is used to initializetime.

As the GPS is acquiring sat-ellites, position, time, andvelocity estimates can bechecked to ensure that theyare within startup error lim-its. If so, monitor STATpage 1. If not, a normal startis required.

After SAT 4 is achieved withgood EPE (Estimated Posi-tion Error), and FM (Figureof Merit) of FM3 or below,check position, velocity, andtime.

GPS is now ready for nor-mal navigation.

(c.) GPS Cold Start.

1. Mode selector switch - INIT.

2. Data selector switch - OPT.

3. Slew key - Select page 4.

4. Enter 04 on line 1.

5. Line select key 2 - Press nextto COLD START to initiate.

6. Line select key 3 - Press nextto COLD START to clearcold start message and re-sume normal display.

7. Mode selector switch - NAV.

e. CHAALS Use of GPS and INS.

(1.) CHAALS Concept. CHAALS (Coher-ent High Accuracy Airborne Location System), is anemitter location system that provides timely, highaccuracy locations required for targeting and to sup-port emitter associations and battlefield situation as-sessment. CHAALS provides this capability throughcoherent processing of differential doppler (DD) andtime difference of arrival (TDOA) information re-ceived at a ground facility from the aircraft.

CHAALS receivers aboard the aircraft will re-ceive and digitize emitter signals. The data will betransmitted over the data link to the GR/CS inte-grated processing facility (IPF). There, CHAALSprocessors will perform the required computationsto produce accurate emitter locations. The precisenavigation required will be provided by the inertialnavigation system (INS) and the global positioningsystem (GPS). GPS also provides the primary means

Change 2 3-76E

TM 55-1510-221-10

of time synchronizing the CHAALS receivers (signalconditioners or SC’s) aboard the aircraft. A backupfor the GPS will be provided by the data link. Theresultant emitter reports will be sent to GR/CS byCHAALS.

(2.) GPS (and INS) Involvement. The ac-curate and timely navigation (position and velocity)is provided by integrating an INS with a GPS, andintegrating both (through a series of intermediaries)with a CHAALS ground based navigation processor(NP). The SC, data link, and CHAALS HSSP (HighSpeed Signal Processor) from the communicationlink. The critical airborne interfaces for CHAALSnavigation and time synchronization include the fol-lowing:

1. INS to GPS:

a. Acceleration

b. Velocity

c. Position

d. Altitude

2. INS to CHAALS: Same as INS to GPS

3. GPS to CHAALS:

a.

b.

c.

d.

e.

Time mark pulse (time synchroniza,tion)

Navigation data block (position, velocity, and time)

Error state vector data block (9 ele-ment ESV, time)

TM/covariance data block (time, TMtime, covariance)

Status data block (status includingDOP’s and FOMN)

3-76F Change 2

TM 55-1510-221-10

Section IV. TRANSPONDER AND RADAR

3-31. WEATHER RADAR SET (AN/APN-215).

a. Description. The weather radar set fig.3-25) provides a visual The weather radar set (fig.3-25), provides a visual 120° around the nose of theaircraft, extending to a distance of 240 nauticalmiles. The presentation on the screen shows thelocation of potentially dangerous areas, such asthunderstorms and hailstorms, in terms of distanceand azimuth with respect to the aircraft. The radaris also capable of ground mapping operations. Theweather radar set is protected by a 5-ampereRADAR circuit breaker located on the overhead cir-cuit breaker panel (fig. 2-26).


(1.) GAIN control. Used to adjust radarreceiver gain in the MAP mode only.

(2.) STAB OFF switch. Push type on/offswitch. Used to control antenna stabilization signals.

(3.) Range switches. Momentary actiontype switches. When pressed, clears the screen andincreases or decreases the range depending on switchpressed.

(4.) TILT control. Varies the elevationangle of radar antenna a maximum of 15 degrees upor down from horizontal attitude of aircraft.

(5.) 60° switch. Push type on/off switch.When activated, reduces antenna scan from 120° to60 degrees.

(6.) TRACK switches. Momentary actiontype switches. When activated, a yellow track lineextending from the apex of the display through toprange mark appears and moves either left or right,depending on the switch pressed. The track lineposition will be displayed in degrees in the upperleft comer of the screen. The line will disappearapproximately 15 seconds after the switch isreleased. It will then automatically return to “0“degrees.

(7 . ) HOLD swi tch . Push type on/offswitch. When activated, the last image presentedbefore pressing the switch is displayed and held. Theword HOLD will flash on and off in the upper leftcomer of the screen. Pressing the switch again willupdate the display and resume normal scan opera-tion.

(8.) Function switch. Controls operation ofthe radar set.


(b.) STBY. Places set in standbymode. This position also initiates a 90-secondwarm-up delay when first turned on.

(c.) TEST. Displays test pattern tocheck for proper operation of the set. The transmit-ter is disabled during this mode.

tion.(d.) ON. Places set in normal opera-

(9.) MODE switches. Momentary actiontype switches. Pressing and holding either switchwill display an information list of operational dataon the screen. The data heading will be in blue, alldata except present data will be in yellow, and pres-ent selected data will show in blue. The threeweather levels will be displayed in red, yellow, andgreen. If WXA mode has been selected, the red barwill flash on and off. If the switch is released andimmediately pressed again, the mode will increaseor decrease depending on switch pressed. Wheneither top or bottom mode is reached, the oppositeswitch must be pressed to further change the mode.

(10.) NAV switch. If pressed with the INSoperating and the weather radar operating in aweather depiction mode, the screen will display INSwaypoints that are located within the range dis-played and within the degree of coverage left or rightof the present heading of the aircraft.

(11.) BRT control. Used to adjust screenbrightness.

c. Weather Radar - Normal Operation.

Do not operate the weather radar setwhile personnel or combustible materialsare within 18 feet of the antenna reflectorWhen the weather radar set is operating,high-power radio frequency energy isemitted from the antenna reflector, whichcan have harmful effects on the humanbody and can ignite combustible materi-als.

3-77

TM 55-1510-221-10

Do not operate the weather radar set in aconfined space where the nearest metalwall is 50 feet or less from the antennareflector. Scanning such surfaces maydamage receiver crystals.

(1.) Turn-on procedure: Function switch -TEST or ON, as required. (Information will appearafter time delay period has elapsed.)

(2.) Initial adjustment operating proce-dure:

1. BRT control - As required.

2 . MODE swi tches - Press andrelease as required.

3. RANGE switches - Press andrelease as required.

4. TILT control - Move up or downto observe targets above or belowaircraft level. The echo displaywill change in shape and locationonly.

(3.) Test procedure:

1. Function switch - TEST.

2 . RANGE swi tches- Press andrelease as required to obtain 80-mile display.


4. Screen - Verify proper display.(The test display consists of twogreen bands, two yellow bands,and a red band on a 120-degreescan. The word TEST will be dis-played in the upper right comer.The operating mode selected bythe MODE switches, either MAP,WX, or WXA, will be displayedin the lower left corner. If WXAhas been selected, the red band inthe test pattern will flash on andoff. The range will be displayed inthe upper right corner beneath theword TEST and appropriate rangemark distances will appear alongthe right edge of the screen.)

(4.) Weather observation operating proce-dure:

1. Function switch - ON.

2.

3.

4.

5.

6.

MODE switches - Press andrelease as required to select WX.

BRT control - As required.

TILT control - Adjust untilweather pattern is displayed.Include the areas above andbelow the rainfall areas to obtaina complete display.

MODE switches - Press andrelease to select WXA. Areas ofintense rainfall will appear asflashing red. These areas must beavoided.

TRACK switches - Press to movetrack line through area of leastweather intensity. Read relativeposition in degrees in upper leftcorner of screen.

NOTE

Refer to FM 1-30 for weather observa-tion, interpretation and application.

(5.) Ground mapping operating procedure:

1. Function switch - ON.

2. MODE swi tches - Press andrelease as required to select MAP.


4. GAIN control - As required topresent usable display.

(6.) Standby procedure: Function switch -STBY.

(7.) Shutdown procedure: Function switch- OFF.

(8.) Weather radar emergency operation.Not applicable.

3-32. TRANSPONDER SET (APX-100).

a. Description. The transponder systemreceives, decodes, and responds to interrogationsfrom Air Traffic Control (ATC) radar to allow air-craft identification, altitude reporting, positiontracking, and emergency tracking. The systemreceives a radar frequency of 1030 MHz and trans-mits preset coded reply pulses on a radar frequencyof 1090 MHz at a minimum peak power of 200watts. The range of the system is limited to line-of-sight.

3-78

TM 55-1510-221-10

Figure 3-26. Weather Radar Control-Indicator (AN/APN-215)

3-79

TM 55-1510-221-10

The transponder system consists of a combinedreceiver/transmitter/ control panel (fig. 3-27) locatedon the pedestal extension; a pair of remote switches,one on each control wheel; and two antennas,located on the underside and top of the fuselage.The system is protected by the 3-ampere TRAN-SPONDER and the 35-ampere AVIONICS MAS-TER PWR No. 1 circuit breakers on the overheadcircuit breaker panel (fig. 2-26).


(1.) TEST-GO indicator. Illuminates toindicate successful completion of built-in-test (BIT).

(2.) TEST-MON indicator. Illuminates toindicate system malfunction or interrogation by aground station.

(3.) ANT switch. Selects desired antennafor signal input.

antennas.

(a.) TOP. Selects upper antenna,

(b.) DIV. Selects d iverse (both)

(c.) BOT. Selects lower antenna.(9.) IDENT-MIC-OUT switch. Selects

source of aircraft indentification signal.

(4.) RAD TEST-OUT switch. Enablesreply to TEST mode interrogations from test set.

(a.) IDENT. Activates transmissionof identification pulse (1P).

(5.) MASTER CONTROL. Selects systemoperating mode.

(a.) OFF. Deactivates system.

(b.) STBY. Activateswarm-up (standby) mode.

system

(c.) NORM. Activates normal oper-ating mode.

(d.) EMER. Transmits emergencyreply code.

(6.) STATUS ANT indicator. Illuminatesto indicate the BIT or MON fault is caused by highVSWR in antenna.

(7.) STATUS KIT indicator. Illuminates toindicate the BIT or MON fault is caused by externalcomputer.

(8.) STATUS ALT indicator. Illuminatesto indicate the BIT or MON fault is caused by thealtitude digitizer.

Figure 3-27. Transponder Control Panel (AN/APX - 100)

3-80

(b.) MIC. Enables either controlwheel POS IDENT switch to activate transmissionof ident signal from transponder.

(c.) OUT. Disallows outgoing signal.

(10.) MODE 4 reply indicator light. Illumi-nates to indicate a reply has been made to a validMode 4 interrogation.

(11.) MODE 4 AUDIO OUT switch. Selectsmonitor mode for mode 4 operation.

(a.) AUDIO. Enables sound and sightmonitoring of mode 4 operation.

(b.) LIGHT. Enables moni tor ingREPLY indicator for mode 4 operation.

mode.(c.) OUT. Deactivates monitor

(12.) MODE 3/A code selectors. Selectdesired reply codes for mode 3/A operation.

(13.) MODE 1 code selectors. Select desiredreply codes for mode 1 operation.

(14.) MODE 4 TEST-ON-OUT switch.Selects test mode of Mode 4 operation.

(a.) TEST. Activates built-in-test ofmode 4 operation.

tion.

(b.) ON. Activates mode 4 operation.

(c.) OUT. Disables mode 4 opera-

(15.) MODE 4 code control. Selects presetmode 4 code.

(16.) M-C, M-3A, M-2, and M-1 switches.Select test or reply mode of respective codes.

(a.) TEST. Activates self-test ofselected code. Transponder can also reply

(b.) ON. Activates normal operation.

(c.) OUT. Deactivates operation ofselected code.

(17.) MODE 2 code selectors. Select desiredreply codes for Mode 2 operation. The cover overmode select switches must be slid forward to displaythe selected mode 2 code.

(18.) POS IDENT pushbutton (controlwheels, jig. 2-17). When pressed, activates transpon-der identification reply.

TM 55-1510-221-10

c. Transponder - Normal Operation.

(1.) Turn-on procedure: MASTER switch -STBY. Depending on the type of receiver installed,the TEST/MON NO GO indicator may illuminate.Disregard this signal.

(2.) Test procedure:

NOTE

Make no checks with the master switch inEMER, or with M-3/A codes 7600 or7700 without first obtaining authorizationfrom the interrogating station(s).

1.

2.

3.

4.

5.

6.

7.

8.

9.

Allow set two minutes to warmup.Select codes assigned for use inmodes 1 and 3/A by depressingand releasing the pushbutton foreach switch until the desirednumber appears in the properwindow.

Lamp indicators - Operate press-to-test feature.

M-1 switch - Hold in TEST.Observe that no indicator lightsilluminate.

M-1 switch - Return to ON.

Repeat steps 4 and 5 for the M-2,M-3/A and M-C mode switches.

MASTER control - NORM.

MODE 4 code control - A. Set acode in the external computer.

MODE 4 AUDIO OUT switch -OUT.

(3.) Modes 1, 2, 3/A, and/or 4 operatingprocedure.

NOTE

If the external security computer is notinstalled, a NO GO light will illuminateany time the Mode 4 switch is moved outof the OFF position.

1. MASTER control - NORM.

2. M-1, M-2, M-3/A, and/or MODE4 ON-OUT switches - ON. Actu-ate only those switches corre-sponding to the required codes.The remaining switches should beleft in the OUT position.

3-81

TM 55-1510-221-10

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

16.

17.

18.

19.

MODE 1 code selectors - Set (ifapplicable).

MODE 3/A code selectors - Set (ifapplicable).

MODE 4 code control - Set (ifrequired).

MODE 4 REPLY indicator -Monitor to determine when tran-sponder set is replying to a SIFinterrogation.

MODE 4 AUDIO OUT switch -Set (as required to monitor Mode4 interrogations and replies).

MODE 4 audio and/or indicator -Listen and/or observe (for Mode 4interrogations and replies).

IDENT-MIC-OUT switch - Pressto IDENT momentarily.

MODE 4 TEST-ON-OUT switch- TEST.

Observe that the TEST GO indi-cator light illuminates.

MODE 4 TEST-ON-OUT switch- ON.

ANT switch - BOT.

Repeat steps 4, 5, and 6. Observethat the TEST GO indicator illu-minates.

TOP-DIV-BOT-ANT switch -TOP.

Repeat step 14.

TOP-DIV-BOT-ANT switch -DIV.

Repeat step 14.

When possible, obtain the cooper-ation of an interrogating stationto exercise the TEST mode. Exe-cute the following steps:

a. RAD TEST-OUT switch -RAD TEST.

b . Obta in ver i f ica t ion f rominterrogating station that aTEST MODE reply wasreceived.

c. RAD TEST-OUT switch -OUT.

(4.) Transponder set identification-position d. Transponder - Emergency Operation. Notoperating procedure: The transponder set can make applicable.

identification-position replies while operating incode Modes 1, 2, and/or 3/A, in response to groundstation interrogations. This type of operation is initi-ated by the operator as follows:

1. Modes 1, 2, and/or 3/A - ON, asrequired.

2. IDENT-OUT-MIC switch - Pressmomentarily to IDENT, whendirected.

NOTE

Holding circuits within the transponderreceiver-transmitter will transmit identifi-cation-position signals for 15 to 30 sec-onds. This is normally sufficient time forground control to identify the aircraft’sposition. During the 15 to 30 secondperiod, it is normal procedure to acknowl-edge via the aircraft communications setthat identification- position signals arebeing generated.

NOTE

Set any of the M1, M2, M3/A, M-C, orMODE 4 switches to OUT to inhibittransmission of replies in undesiredmodes.

NOTE

With the IDENT-OUT-MIC switch set tothe MIC position, the POS IDENT but-ton must be depressed to transmit identi-fication pulses.


1. To retain Mode 4 code in externalcomputer during a temporaryshutdown:

a. MODE 4 CODE switch -Rotate to HOLD.

b. Wait 15 seconds.

c. MASTER control - OFF.

2. To zeroize the Mode 4 code in theexternal computer turn MODE 4CODE switch to ZERO.

3. MASTER control - OFF. This willautomatically zeroize the externalcomputer unless codes have beenretained (step 1 above).

3-82

TM 55-1510-221-10

Figure 3-28. Pilot’s Altimeter Indicator (BA-141)

3-33. PILOT’S ALTIMETER INDICATOR.

The pilot’s altimeter, on the upper left side ofthe instrument panel (fig, 3-28), is a servoed unitunder control of the Air Data Computer and is partof the Flight Director/Autopilot system. Altitude isdisplayed by a 10,000 foot counter, a 1000 footcounter, a 100 foot counter, and a single needlepointer (coupled with the 100 foot counter) whichindicates hundreds of feet on a circular scale in 20foot increments. Below an altitude of 10,000 feet, adiagonal striped symbol will appear on the 10,000foot counter. The barometric pressure knob allowsground supplied pressure values to be adjusted anddisplayed in inches Hg or millibars. If AC power tothe altimeter is lost, a warning OFF flag will appearin the upper counter drum display window to indi-cate power loss, unreliable altimeter readings, andpossible loss of encoder transmissions to ground sta-tions. Circuits are protected by a 3-ampere fuse in ajunction box.

When the BARO knob is adjusted to groundsupplied instructions, the updated altitude pressureis routed to the Air Data Computer. The ADCrecomputes all data on hand, sends corrected alti-tude pressure information to the Flight Director andautopilot, servo commands to correct the display onthe pilot’s altimeter, and supplies altitude informa-

tion to the transponder (for transmission to .

a. Controls/Indicators and Functions.

(1.) ALT alert annunciator. Illuminateswhen aircraft is within 1000 feet of preselected alti-tude during capture maneuver and extinguisheswhen aircraft is within 250 feet of preselected alti-tude. After capture, light will illuminate if aircraftdeparts more than 250 feet from the selected alti-tude.

(2.) Altitude counter drums. Indicates air-craft altitude in tens of thousands, thousands, andhundreds of feet above sea level.

(3.) IN HG Indicator. Indicates local baro-metric pressure in inches of mercury. Adjusted byuse of BARO knob.

(4.) Needle indicator. Indicates aircraftaltitude in hundreds of feet with subdivisions at 20foot increments.

(5.) BARO Knob. Used to manually setbarometric pressure displayed in the MB and INHG windows.

3-83

TM 55-1510-221-10

(6.) MB Indicator. Indicates local baro-metric pressure in millibars. Adjusted by use ofBARO knob.

(7.) Below 10,000 feet symbol. Presenceindicates aircraft altitude is below 10,000 feet.

(8.) OFF Flag. Presence indicates loss ofpower to instrument and unreliable readings.

NOTE

If the OFF flag is visible, either DCpower is off, the fuse has blown, or thereis an altimeter encoder failure. Since theOFF flag monitors only the encoder inputto the altimeter and not transponder con-dition, the altitude reporting functionmay be inoperative without the OFF flagshowing, in the case of transponder fail-ure or improper control settings. It is alsopossible to get a good Mode C test on thetransponder control with the OFF flagshowing. If the OFF flag remains visible,radio contract should be made with aground radar site to determine if the alti-tude reporting function is operative.

b. Pilot’s Altimeter - Normal Operation.

(1.) Turn-on procedure: Servoed altimeterwill operate when transponder is operating withM-C switch set to center position.

(2.) Operating procedure:

1. Barometric set knob - Set desiredaltimeter setting in IN. HG. win-dow.

2. OFF flag - Check not visible.

3. Needle indicator - Check opera-tion.

NOTE

If the altimeter does not read within 70feet of field elevation, when the correctlocal barometric setting is used, the altim-eter needs calibration or internal failurehas occurred. An error of greater than 70feet also nullities use of the altimeter forIFR flight.

c. Pilot’s Altimeter - Emergency Operation.Disregard pilot’s altimeter and utilize copilot’saltimeter.

3-34. COPILOT’S ENCODING ALTIMETER.

Description. The copilot’s altimeter (fig.3-29):’ provides an indication of present aircraftpressure altitude above sea level. It also suppliesinformation to the INS and GPS. The air data com-puter supplies altitude information to the transpon-der.


(1.) ALT alert indicator. Not used.

(2.) Needle indicator. Indicates aircraftaltitude in hundreds of feet with subdivisions at 20-foot intervals.

(3.) MILLIBARS window. Indicates localbarometric pressure in millibars. Adjusted by use ofset knob.

(4.) IN HG window. Indicates local baro-metric pressure in inches of mercury. Adjusted byuse of set knob.

(5.) BARO knob. Used to manually setbarometric pressure displayed in the MB an IN HGwindows.

(6.) Drum indicator. Indicates aircraft alti-tude in ten-thousands, thousands, and hundreds offeet above sea level.

(7.) Test button. Used to test altimeteroperation.

c. Encoding Altimeter - Normal Operation.

(1.) Turn-on procedure: Encoding altime-ter will operate when transponder is operating withM-C switch set to center position.

(2.) Operating procedure:

1. Barometric set knob - Set desiredaltimeter setting in IN. HG. win-dow.

2. CODE OFF flag - Check not visi-ble.

3. Needle indicator - Check opera-tion.

4. TEST button -

a. Push - Reading decreases by500 feet.

b. Release - Returns to originalreading.

3-84

TM 55-1510-221-10

1. Altitude alert annunciator2. Altitude pointer3. MILLIBARS barometric pressure counter4. IN HG barometric pressure counter5. BARO knob6. TEST button7. Counter drum display

AP 012930

Figure 3-29. Copilot’s Encoding Altimeter

NOTE feet also nullifies use of the altimeter for

If the altimeter does not read within 70feet of field elevation, when the correctlocal barometric setting is used, the altim-eter needs calibration or internal failurehas occurred. An error of greater than 70

IFR flight.

d. Encoding Altimeter - Emergency Operation.Altimeter circuit breaker - Pull (if encoder faultoccurs).

3-85/(3-86 blank)

TM 55-1510-221-10

CHAPTER 4

Mission Equipment

Section I. MISSION AVIONICS

4-1. MISSION AVIONICS COVERAGE.

Complete provisions only are installed for theGPS, CHAALS and AQL mission systems. Equip-ment descriptions and operating instructions are tobe obtained from appropriate vendor and ArmyTechnical manuals.

4-2. MISSION CONTROL PANEL.

The mission control panel (fig 4-l), mounted onthe copilot’s sidewall, consists of three sections. The

top section contains the mission caution/advisoryannunciator panel, see Table 2-8. The center sectioncontains one DC volt/ammeter, two digital AC volt/frequency meters, two AC digital load meters, oneantenna steering synchro control, and the antennasteering mode selector switch. The bottom sectioncontains the mission equipment control switchesand the mission equipment circuit breakers.

Section II. AIRCRAFT SURVIVABILITY EQUIPMENT

4-3. M-130 FLARE AND CHAFF DISPENSINGSYSTEM.

a. Description. The M- 130 flare and chaff dis-pensing system provides effective survival counter-measures against radar guided weapons systems andinfrared seeking missile threats. The system consistsof two dispenser assemblies with payload moduleassemblies, a dispenser control panel, a flare dis-pense switch, two control wheel mounted chaff dis-pensing switches, an electronic module assembly,and associated wiring. The flare and chaff dispens-ing system is protected by a 5-ampere circuitbreaker, placarded M130 POWER located on themission control panel (fig. 4-l).

Right engine nacelle dispenser is for chaffonly.

(1.) Dispenser assemblies. Two inter-changeable dispenser assemblies are mounted on theaircraft. One is located in the aft portion of the rightnacelle and the other is mounted on the right side ofthe fuselage. On this aircraft the dispenser in thenacelle will be used for chaff only while the dis-penser mounted on the fuselage can be used foreither flares or chaff. The selector switch (placardedC-F) on the dispenser can be set for either chaff orflares. The unit also contains the sensor for the flare

detector. The dispenser assembly breech plate hasthe electrical contact pins which fire the impulse car-tridges. The unit also contains the sequencing mech-anism.

(2.) Payload module assemblies. A remov-able payload module assembly is provided for eachdispenser assembly. Each payload module has 30chambers which will accept either flares or chaffs.Flares or chaffs are loaded into the rear-end (stud-ded end) of the payload module, and secured inplace by a retaining plate.

(3.) Electronic module assembly (EM).The electronic module assembly contains the pro-grammer, the flare detector and a safety switch. Theunit is located behind the pilot’s seat.

(a.) Flare detector. The flare detectoris provided to insure that a flare is burning when itis ejected from the dispenser payload module. If theinitial flare fails to ignite, the detector automaticallyfires another flare within 75 milliseconds. If the sec-ond flare fails to ignite, the detector will fire a thirdflare. If the third flare ignition is not detected, thedetector will not fire another flare until the systemis activated again by pressing the FLARE DIS-PENSE switch.

(b.) Programmer. The programmer isused for the chaff mode only. It has four switchesfor setting count and interval of salvo and burst.

(c.) Safety switch. The safety switch(with safety pin and red flag) prevents firing of chaff

4-1

TM 55-1510-221-10

Figure 4- 1. Mission Control Panel

4-2

TM 55-1510-221-10

or flares when the safety pin is inserted. The safetypin shall be removed only while the aircraft is inflight or during test of the system.

(4.) Flare dispenser switch. A single push-button switch (fig. 4-2) placarded FLARE DIS-PENSE, located on the control pedestal, will fire aflare from the dispenser payload module each timeit is pressed. If the FLARE DISPENSE switch isheld down, it will dispense a flare every 2.3 seconds.

(5.) Control wheel mounted chaff dispenseswitches. Two pushbutton switches placardedCHAFF DISP, one located on top left portion of thepilot’s control wheel and the other located on thetop right portion of the copilot’s control wheel, acti-vates the chaff dispensing 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 thenacelle, prevents the firing of chaff or flares whenthe pin is inserted. This safety pin shall be insertedwhile the aircraft is on the ground and removedprior to flight or during system test.

(7.) Dispenser control panel (DCP). Theflare dispenser control panel (fig. 4-3) is mounted inthe control pedestal. Control functions are as fol-lows:

(a.) RIPPLE FIRE switch. A guardedswitch placarded RIPPLE FIRE fires all remainingflares when moved to the up position. It is used inthe event of an inflight emergency to dispense allflares from the dispenser payload module.

(b.) FLARE counter setting knob.Facilitates setting FLARE counter to the number offlares in the payload module before flight.

(c.) FLARE counter. Indicates thenumber of flares remaining in the dispenser payloadmodule.

(d.) ARM light. An amber press totest indicator light placarded ARM illuminates whenthe ARM-SAFE switch is in the ARM position,when the safety pins are removed from the elec-tronic module and the wing safety switch. Clockwiserotation will dim the indicator light.

(e.) CHAFF counter. Indicates thenumber of chaffs remaining in the payload module.

(f) CHAFF counter setting knob.Facilitates setting CHAFF counter to the number ofchaffs in the payload module before flight.

(g.) MAN-PGRM SELECTORSWITCH. Selects manual or programmed chaff dis-pense.

AP006537

Figure 4-2. Flare Dispense Switch

4-3

TM 55-1510-221-10

(h.) ARM-SAFE switch. When in theSAFE position, power is removed from the M-130

(2.)

system. When in the ARM position, power isignated crew

applied to the M-130 system.the safety pinfor replacing

1. MAN. Bypasses the pro-grammer and fires one chaff each time one of thechaff dispense switches is pressed.

2. PGRM. Chaff is fired inaccordance with the preset chaff program as set intothe electronic module (count and interval of burstsand salvo).

(i.) Ripple Fire Switch Cover. Pre-vents accidental switch activation.

aircraft is airborne the pilot is responsible forremoving the safety pin from the electronic moduleand moving the ARM-SAFE switch on the dispensercontrol panel to ARM. Before landing, he is respon-sible for re-inserting the safety pin in the electronicmodule and moving the ARM-SAFE switch toSAFE. While airborne the pilot and copilot areresponsible for scanning the terrain for missilethreats. When either pilot recognizes a missilelaunch he will press the FLARE DISPENSE buttonto eject flares.

(8.) Ammunition for dispenser. Ammuni-tion for the system consists of countermeasure chaffMl, countermeasure flares M206, and impulse car-tridges M796.

(a.) Countermeasure chaff M1. Theseunits consist of a plastic case 8 inches in length and0.97 inches square. The base of the chaff case isflanged to provide one-way assembly into the dis-penser payload module. The chaff consists of alumi-num coated fiberglass strands.

(3.) Conditions for firing. The dispensersystem should not be fired unless a missile launch isobserved or radar guided weapons systems isdetected and locked on. If a system malfunction issuspected, aircraft commander may authorizeattempts to dispense flares or chaff as a test in anon-hostile area.

(b) Countermeasure flare M206.These units consist of an aluminum case 8 inches inlength and 0.97 inches square. The base of the flareis flanged to provide one-way assembly into the pay-load module. The flare material consists of a magne-sium and teflon composition. A preformed packingis required in the base of the flare unit prior toinserting the impulse cartridge.

Aircraft must be in flightflares.

(c.) Impulse cartridge M796. Thiscartridge fits into the base of either the flare or chaffand is electrically initiated to eject flares or chafffrom the dispenser payload module.

b. Normal Operation.

1. Flares. Upon observing amissile launch the pilot or copilot (whoever sightsthe launch first) will fire a flare. If more than onemissile launch is observed, the firing sequenceshould be continued until the aircraft has cleared thethreat area.

NOTE

If aircraft is to be flown with flare dis-penser assembly removed, fairing shouldbe removed from fuselage.

(1.) General. At the present time surface-to-air intermediate range guided missiles that arelaunched against the aircraft must be visuallydetected by the aircraft crew. Crew members mustinsure visual coverage over the ground area where amissile attack is possible. The aircraft radar warningsystem will only alert the pilot and copilot when theaircraft is being tracked by radar-guided anti-aircraftweapons systems. It will not indicate the tiring ofweapons against the aircraft.

2_ Chaff Upon receiving analert from the aircraft radar warning system, thepilot or copilot will fire the chaff and initiate an eva-sive maneuver. The number of burst/salvo and num-ber of salvo/program and their intervals as estab-lished by training doctrine will be set into theprogrammer prior to take-off (refer to TM 9-1095206- 13 & P for information on setting programmer).If desired, the operator may override the pro-grammed operational mode and fire chaff counter-measures manually by moving the dispenser func-tion selector switch to MANUAL and pressing adispenser switch.

(b.) Firing responsibility. When thepilot or copilot observes a missile launch or radarwarning indication, he fires flares or chaff andassumes command of the dispenser system, and firessucceeding flares as required. He will advise theother pilot that a missile launch has been observed

4-4

Crew responsibilities. The pilot or des-member is responsible for removingfrom the right wing before flight, andit immediately after flight. After the

to dispense

(a.) Firing procedure.

TM 55-1510-221-10

1. RIPPLE FIRE switch2. FLARE counter setting knob3. FLARE counter indicator4. ARM light5. CHAFF counter indicator6. CHAFF counter setting knob7. MANUAL-PROGRAM switch8. ARM-SAFE switch9 RIPPLE FIRE switch cover

Figure 4-3. Flare Dispenser Control Panel

AP 006768

4-5

TM 55-1510-221-10

or a radar warning signal has been received, andannounce that flares or chaff have been fired.

4-4. SYSTEM DAILY PREFLIGHT/RE-ARM TEST.

The following test procedures shall be conductedprior to the first flight of each day and prior to eachre-arming of the dispensers. The first dispensertested shall be the one used to dispense flares andthe second one shall be the one used to dispensechaff. Notify AVUM if any improper indicationsoccur during the tests.

Assure payload module is not connectedto dispenser assembly at any time duringthe following test procedure.

1. On flare dispenser assembly, assure the C-Fselector switch is in F (flare) position.

2. Obtain M-91 test set and assure 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 inchhexagonal wrench provided in test set carry-ing case.

4. Obtain test set power cable from the M-91test set carrying case and connect cablebetween exterior connection J 1 (28V DC) onaircraft and aircraft power + 28V DC (J1)of test set.

5. Remove safety pin from EM and in the topskin of the right wing.

On DCP, assure that RIPPLE FIREswitch guard is in down position.

6. Provide aircraft power to DCP by settingM- 130 POWER circuit breaker to ON posi-tion.

7. On DCP, press ARM lamp. Lamp will illu-minate. Release ARM lamp. Lamp willextinguish.

8. On DCP, set FLARE counter to 30 CHAFFCOUNTER to 30 and MAN-PGRM switchto MAN position.

9. On DCP, set ARM-SAFE switch to ARM.ARM lamp will illuminate.

NOTE

When the test set is installed on the dis-penser assembly and 28 volts DC aircraftpower has been applied, the sequencerswitch inside of dispenser assembly resets,making an audible sound as it rotates.There will be no such sound if thesequencer switch has been previouslyreset or if switch is in position 12 or 24.

NOTE

On test set, TS PWR ON lamp (clear)illuminates and remains illuminatedthroughout the test sequence until aircraftpower to test set (via test set power cable)is disconnected or shut off.

10. Set mission chaff program on EM.

11. Perform the following operations on theM-91 test set:

a. Press to test the remaining three lampson test set. Each lamp will illuminate.

NOTE

Replace any lamp that does not illumi-nate when pressed. If none of the indicat-ing lamps illuminate, return test set toAVUM.

b. Rota te TEST SEQUENCE swi tchclockwise to the next position, TSRESET. No visual indication willoccur.

c . Rota te TEST SEQUENCE swi tchclockwise to SV SELF TEST position.STRAY VOLTAGE lamp (red) willilluminate.

d . Rota te TEST SEQUENCE swi tchclockwise to next position, TS RESET.STRAY VOLTAGE lamp (red) willextinguish.

e. Rotate TEST SEQUENCE switchclockwise to next position, STRAYVOLT. STRAY VOLTAGE lamp (red)should not illuminate.

f . Rota te TEST SEQUENCE swi tchclockwise to next position, SYS NOTRESET. SYS NOT RESET lamp(amber) should not illuminate. If lamp

4-6

TM 55-1510-221-10

illuminates, press and release MAN-UAL SYSTEM RESET switch and SYSNOT RESET lamp should then extin-guish.

NOTE

When the MANUAL SYSTEM RESETswitch is pressed and released, and 28volts DC power has been applied, thesequencer switch inside the dispenserassembly resets, making an audible soundas it rotates. If the sequencer switch hasbeen previously reset or if the switch is inposition 12 or 24, there will be no suchsound.

g. Rotate TEST SEQUENCE switchclockwise to next position, DISPCOMP.

12. Press FLARE DISP switch once. For eachdepressing, the FLARE counter on DCPshould count down in groups of three.

13. On DCP, raise RIPPLE FIRE switch guardand set toggle switch to up position untilFLARE counter counts down to 00. Returnswitch guard to down position. On DCP,reset FLARE counter back to 30. DIS-PENSER COMPLETE lamp (green) on testset will illuminate.


a. Rotate TEST SEQUENCE switchcounter-clockwise to SYS NOT RESETposition. SYS NOT RESET lamp(amber) will illuminate. DISPENSERCOMPLETE lamp (green) will remainilluminated.

b. Press and release MANUAL SYSTEMRESET switch. SYS NOT RESET lamp(amber) will extinguish.

NOTE


c. Rotate TEST SEQUENCE switchcounterclockwise to STRAY VOLTposition. STRAY VOLTAGE lamp(red) should not illuminate.

d . Rota te TEST SEQUENCE swi tchcounterclockwise to START/HOMEposition.

NOTE

When the TEST SEQUENCE switch isturned to the START/HOME position,the DISPENSER COMPLETE lamp willextinguish, the STRAY VOLTAGE lampwill illuminate and then will extinguishwhen passing through the TS RESETposition.

15. On CHAFF dispenser assembly, assure that C-F selector 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 usingball hexagonal key screwdriver provided intest set carrying case.

NOTE

When the test set is installed on the dis-penser assembly and 28 volts DC aircraftpower has been applied, the sequenceswitch inside the dispenser assemblyresets, making an audible sound as itrotates. There will be no such sound if thesequencer switch has been previouslyreset or if switch is in position 12 or 24.

NOTE

On test set, TS PWR ON lamp (clear)illuminates and remains illuminatedthrough the test sequence until aircraftpower to test set (via test set power cable)is disconnected or shut off.


a. Press to test all four lamps on test set.Each lamp will illuminate.

4-7

TM 55-1510-221-10

NOTE

Replace any lamp that does not illumi-nate when pressed. If none of the indicat-ing lamps illuminate, return test set toAVUM.

b. Rota te TEST SEQUENCE swi tchclockwise to TS RESET position. Novisual indication will occur.

c. Rotate TEST SEQUENCE switchclockwise to SV SELF TEST position.STRAY VOLTAGE lamp (red) willilluminate.

d . Rota te TEST SEQUENCE swi tchclockwise to next position, TS RESET.STRAY VOLTAGE lamp (red) willextinguish.

e. Rotate TEST SEQUENCE switchclockwise to next position, STRAYVOLT. STRAY VOLTAGE lamp (red)should not illuminate.

f . Rota te TEST SEQUENCE swi tchclockwise to next position, SYS NOTRESET. SYS NOT RESET lamp(amber) should not illuminate. If lampilluminates, press and release MAN-UAL SYSTEM RESET switch and SYSNOT RESET lamp should then extin-guish.

NOTE


g . Rota te TEST SEQUENCE swi tchclockwise to next position, DISPCOMPL.

19. Press pilot CHAFF DISP switch once, Presscopilot CHAFF DISP switch once. On DCP,for each depressing, the CHAFF countershould count down by an increment of one.

20. On DCP, set MAN-PGRM switch to PGRMposition.

21. Press any one of CHAFF DISP switches inaircraft. On DCP, the number shown on

22.

23.

CHAFF counter should decrease in accor-dance with the program set on the EM.

Repeatedly press other CHAFF DISPENSEswitch until CHAFF counter on DCP reads00.

On test set, observe DISPENSE COM-PLETE lamp (green) is illuminated and thenperform the following operations:

a. Rotate TEST SEQUENCE switchcounter-clockwise to SYS NOT RESETposition. SYS NOT RESET lamp(amber) will illuminate.

b. Press and release MANUAL SYSTEMRESET switch. SYS NOT RESET lamp(amber) will extinguish.

NOTE


c. Rotate TEST SEQUENCE switchcounter-clockwise to STRAY VOLTposition. STRAY VOLTAGE lamp(red) should not illuminate.

d . Rota te TEST SEQUENCE swi tchcounter-clockwise to START/HOMEposition.

NOTE

When the TEST SEQUENCE switch isturned to the OFF position, the DIS-PENSER COMPLETE lamp will extin-guish, the STRAY VOLTAGE lamp willilluminate and then will extinguish whenthe OFF position is reached.

24. Install safety pins.

25. Disconnect test set power cable.

26. Remove M-91 test set from dispenser assem-bly and restore in carrying case along withthe power cable and hexagonal wrench.

27. On DCP, set ARM-SAFE switch to SAFEposition.

28. On DCP, reset CHAFF counter to 30.

4-8

TM 55-1510-221-10

29. Disconnect aircraft power by pulling the 5ampere M130 POWER circuit breakerlocated on the mission control panel (fig.4-l).

30. Proceed immediately to ammunition loadingprocedures.

The system must have been tested toassure that there is no stray voltage andall aircraft power must be removed fromthe system prior to unloading the payloadmodule.

4-5. AMMUNITION.

a. Ammunition Loading Procedure.

Only one shipping container is to beopened at a time. If a shipping containerhas been opened and only partially emp-tied, the remaining contents will besecured in the container with an appropri-ate type of packaging material or filler toadequately prevent jostling. All munitionsin storage must be in their original ship-ping containers.

7. On the dispenser control panel, assureARM-SAFE switch is in SAFE posi-tion.

8. On the electronic module and rightwing assure safety pins and flag assem-blies are installed.

9. Slide payload module assembly intodispenser assembly and secure two studbolts, hand tight, using 5/32 inch hex-agonal wrench.

b. Ammunition Unloading Procedure.

1. Place payload module assembly onwork bench in approved safe area sothat the retaining plate is facing up.

All aircraft power to the dispenser systemmust be turned off prior to removal ofpayload module from dispenser assembly.Safety pin flag shall be installed in the

2. Remove retaining plate by unscrewing electronic module prior to landing and

two retaining bolts. the safety pin flag shall be installed in thewing-mounted safety switch immediately

3. Insert one flare (or chaff) at a time into after landing.each chamber of payload module.

4. Remove plastic dust cap from eachchaff or flare.

1.

2.

Prior to insertion of an impulse cartridge,be sure there is preformed packing in theflare cartridge. (There will be no pre-formed packing in chaff cartridges.) Rein-stall any preformed packing that is inad-vertently removed with dustcap. Theloading of impulse cartridges into a flareor chaff shall be accomplished one at atime.

If there is an indication that a misfireoccurred, notify emergency ordnance dis-posal personnel for disposition and dis-posal.

3. Remove module from dispenser assem-bly by unscrewing two stud bolts witha 5/32 inch hexagonal wrench and slid-ing out of dispenser assembly.

5. Insert one impulse cartridge into eachflare (or chaff).

6 . Install retainer plate assembly by 4. Remove retaining plate from payloadscrewing to two retainer bolts into pay- module by unscrewing two retainingload module. bolts.

On dispenser control panel, assureARM-SAFE switch is in SAFE posi-tion.

Assure safety pin and flag are insertedinto electronic module and in the wingmounted safety switch.

4-9

TM 55-1510-221-10

5. Remove expended and unexpendedimpulse cartridges and flares (or chaff)from payload module.

6. Repack unexpended items in originalcontainers and return to stores.

NOTE

It is not unusual for the case of a chaffcartridge to crack when fired. It does noteffect performance of the item and shouldnot be reported as a malfunction.

4-6. RADAR SIGNAL DETECTING SET (AN/APR-39(V)1).

The radar signal detecting set (control panel, fig.4-3) indicates the relative position of search radarstations. Audio warning signals are applied to thepilot’s and copilot’s headsets. The radar signaldetecting set is protected by the 7.5-ampere circuitbreaker placarded APR39, located on the missioncontrol panel (fig. 4-l). The associated antennas areshown in figure 2-1. For operating instructions, referto TM I l-5841-283-20. Pattern # 1 self test, shall beas shown in figure 4-4.

a. Radar Signal Detecting Set Control PanelFunctions (AN/APR-39(V)1) (fig. 4-4).

(1.) PWR switch. Turns set on or off.

(2.) SELF TEST switch. Initiates self test.

(3.) DSCRM switch. Turns discriminatefunction on or off.

(4.) AUDIO control. Adjusts audio level.

b. Radar Signal Detecting Set Indicator Func-tions (fig. 4-5).

(1.) MA indicator. Illuminates to indicatethe presence of an MA threat.

(2.) Display. Indicates relative position ofsearch radar stations.

(3.) BRIL control. Adjusts brilliance,

(4.) DA Y-NIGHT control. Rotate to adjustintensity of display.

4-7. RADAR WARNING RECEIVER (AN/APR-44()(V3).

The radar warning receiver (fig. 4-6) indicatesthe presence of certain types of search radar signals.

1. POWER switch2. SELF TEST switch3. DESCRIMINATE function switch4. AUDIO level control

AP 003891

Figure 4-4. Radar Signal Detecting Set Control Panel (AN/APR-39(V)1)

4-10

TM 55-1510-221-10

1. MA indicator2. Display3. BRIL control4. DAY-NIGHT control

AP 005715

Figure 4-5. Radar Signal Detecting Set Indicator

1. Radar warning indicator2. VOLUME control3. POWER switch

AP 003892

Figure 4-6. Radar Warning Receiver Control Panel (AN/APR-44() (V3)

4-11

TM 55-1510-221-10

The radar warning receiver is protected by the 5-am-pere circuit breaker placarded APR44, located onthe mission control panel (fig. 4-1). For operatinginstructions, refer to TM 11-5841-291-12.

a. Radar warning indicator. Illuminates toindicate the presence of an AI or SAM threat.

b. VOLUME control. Adjusts volume.

c. POWER switch. Turns set on and off.

4-12

TM 55-1510-221-10

5-1. PURPOSE.

CHAPTER 5Operating Limits and Restrictions

Section I. GENERAL

This chapter identifies or refers to all importantoperating limits and restrictions that shall beobserved during ground 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 limitswill allow the pilot to safely perform the assignedmissions and to derive maximum utility from theaircraft. Limits concerning maneuvers, weight, andcenter of gravity are also 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 aircraft operationis two pilots. Additional crewmembers as required willbe added at the discretion of the commander, inaccordance with pertinent Department of the Armyregulations.

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 coding ofeach 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 55-1510-221-10

TORQUE 49% MAXIMUM BELOW 1600 RPM

2O-100% NORMAL OPERATING RANGE

100% MAXIMUM123% TRANSIENT (5 SECONDS)

TURBINE TACHOMETER (N1 SPEED)52% MINIMUM LOW IDLE88% MAXIMUM REVERSE (ONE MINUTE)101.5% MAXIMUM100.1% MAXIMUM BELOW -48 C102.6% TRANSIENT (10 SECONDS)

TURBINE GAS TEMPERATURE

400-75OºC NORMAL OPERATING RANGE66OºC MAXIMUM LOW IDLE750°C MAXIMUM CONTINUOUS750°C MAXIMUM REVERSE (1 MINUTE)850°C MAXIMUM TRANSIENT1000ºC MAXIMUM STARTING (5 SECONDS)

PROPELLER TACHOMETER1600-2000 RPM NORMAL OPERATING RANGE1900 RPM MAXIMUM REVERSE (1 MINUTE)2000 RPM MAXIMUM2200 RPM TRANSIENT (5 SECONDS)

OIL TEMPERATURE AND PRESSUREOIL TEMPERATURE SCALE

10-99ºC NORMAL OPERATING RANGE0-99ºC CRUISE CLIMB AND RECOMMENDED SPEED-40-99ºC STARTING. LOW IDLE. HIGH IDLE

99ºC MAXIMUM104 C TRANSIENT (5 MINUTES)

OIL PRESSURE SCALE

60 PSI MINIMUM60 TO 85 PSI. 49% TORQUE MAXIMUM

85-135 PSI NORMAL OPERATING ABOVE 21.000 FEET85-105 PSI CAUTION RANGE BELOW 21.000 FEET

105-135 PSI NORMAL OPERATING BELOW 21.000 FEET

200 PSI MAXIMUM STARTING WITH COLD OIL

NOTE: + 10 PSI FLUCTUATIONS ARE ACCEPTABLE

Figure 5-1 Instrument Markings (Sheet 1 of 3)

5-2 Change 4

TM 55-1510-221-10

243 KIAS MAXIMUM (Vmo) (.47 MACH)NOTE

MAXIMUM ALLOWABLE AIRSPEED (RED)STRIPED) POINTER IS SELF ADJUSTING

WITH ALTITUDE

91 KIAS MINIMUM SINGLE-ENGINECONTROL SPEED (Vmca)

127 KIAS ONE-ENGINE INOPERATIVEBEST RATE-OF-CLIMB (Vyse)

78-153 KIAS FULL FLAP OPERATING RANGE

198 KIAS MAXIMUM APPROACH FLAPEXTENSION SPEED

PNEUMATIC PRESSURE

12-20 PSI NORMAL OPERATING RANGE

20 PSI MAXIMUM

PROPELLER DEICER AMMETER

14-18 AMPERES NORMAL OPERATION

RBG

APO12879

Figure 5-1. Instrument Markings (Sheet 2 of 3)

5-3

TM 55-1510-221-10

FUEL QUANTITY

O-265 LBS NO TAKEOFF RANGE

CABIN ALTIMETER AND DIFFERENTIAL PRESSURE

O-6.1 PSI NORMAL RANGE

6.1 PSI MAXIMUM

FLAP POSITION INDICATOR

40% TAKEOFF AND APPROACH

RYG

APOO4768.3Figure 5-1. Instrument Markings (Sheet 3 of 3)

5-4

TM 55-1510-221-10

5-9. AUTOPILOT LIMITATIONS.

WARNING

The RC-12H aircraft is certified with wingtippods installed. Should the pods be removed,the autopilot system must be replaced with astandard C-12D autopilot. Affected wiring mustalso be changed.

b. Fuel Management. Auxiliary tanks will not befilled for flight unless the main tanks are full. Maximumallowable fuel imbalance is 400 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.

a. An autopilot preflight check must be. conductedand 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. Fuel System Anti-Icing. Icing inhibitorconforming to MIL-I-27686 will be added tocommercial fuel, not containing an icing inhibitor,during fueling operations, regardless of ambienttemperatures. The additive provides anti-icingprotection and also functions as a biocide to killmicrobial growth in aircraft fuel systems.

5-11. BRAKE DEICE LIMITATIONS.

c. Operation of the autopilot and yaw damper isprohibited during takeoff and landing, and below 200

The following limitations apply to the brake deice

feet above terrain. Maximum speed for autopilotsystem:

operation is 243 KIAS/O.47 Mach.a. The brake deice system shall not be operated at

d. During a coupled ILS approach do not operateambient temperatures above 15°C.

the propellers in the 1750 to 1850 RPM range.b. The brake deice system shall not be operated

longer 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.

5-10. FUEL SYSTEM LIMITS.

NOTE

Aviation gasoline (AVGAS) contains a form oflead which has an accumulative adverse effecton gas turbine engines. The lowest octaneAVGAS available (less lead content) should beused. If any AVGAS is used the total operatingtime must be entered on DA Form 2408-13.

a. Operating Limits. Operation with FUELPRESS light on is limited to IO hours. Log FUELPRESS light on time on DA Form 2408-13. Onestandby boost pump may be inoperative for takeoff.(Crossfeed fuel will not be available from the side withthe inoperative standby boost pump.) Operation onaviation gasoline is time limited to 150 hours betweenengine overhaul and altitude limited to 20,000 feet withone standby boost pump inoperative. Crossfeedcapability is required for climb, when using aviation

c. Maintain 85% N1 or higher dur ingsimultaneous operation of the brake deice and surfacedeice systems. If adequate pneumatic pressure cannotbe provided for simultaneous operation of the brakedeice and surface deice systems, turn OFF the brakedeice 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.

gasoline above 20,000 feet.

Change 4 5-5

TM 55-1510-221-10

5-13. ENGINE LIMITATIONS.

Observe the following limitations (table 5-1) duringoperation of this aircraft equipped with two Pratt andWhitney of Canada, Ltd. PT6A-41 engines. Eachcolumn is a separate limitation. The limits presented donot necessarily occur simultaneously. Wheneveroperating limits are exceeded, pilot should record thevalue and duration of the condition encountered in theaircraft log. Operation of the engines is monitored byinstruments, with the operating limits marked on theface of each instrument.

CAUTION

Engine operation using only the engine drivenfuel pump without boost pump fuel pressure islimited to 10 cumulative hours. All time in thiscategory shall be entered on DA Form 2408-13for the attention of maintenance personnel.

CAUTION

Use of aviation gasoline is time-limited to 150hours of operation during any Time-Between-Overhaul (TBO) period. It may beused in any quantity with primary or alternatefuel.

5-14. OVERTEMPERATURE AND OVERSPEEDLIMlTATlONS.

a. Whenever the limiting temperatures areexceeded and cannot be controlled by retarding thepower levers, the engine will be shut down and alanding made as soon as possible.

b. During engine operation, the temperatures,speeds and time limits listed in the Engine OperatingSpeeds Limitations chart (table 5-1) must be observed.When these limits are exceeded, the incident will beentered as 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 725ºC willreduce engine life.

5-6

5-15. POWER DEFlNlTlONS FOR ENGINEOPERATIONS.

The following definitions describe the enginepower ratings.

a. Takeoff Power. The maximum power avail-able, 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.

5-16. GENERATOR LlMlTS.

Maximum generator load is limited to 100% forflight and variable during ground operations. Observethe limits shown in Table 5-2 during ground operation.

Section ill. POWER LlMlTS

TM 55-1510-221-10

Table 5-1. Operating Limits

OPERATINGCONDITION

GASTORQUE MAXIMUM GENERATOR PROP OIL OIL

SHP PERCENT OBSERVED RPM N1 (10) RPM PRESS TEMP(1) TGT ºC RPM % N2 PSI (2) ºC

TAKEOFF (3) 850 100% 750 38,100 101.5 2000 105 to 10 to 99135

MAX CONT 850 100% (4) 750 38,100 101.5 2000 105 to 10 to 99135

MAX CLIMB 850 100% (4) 725 38,100 101.5 2000 105 to 0 to 99135

MAX CRUISEHIGH IDLE - - - - (5) - - -40 to 99LOW IDLE - - 660 (6) 19,500 52(min) - 60(min) -40 to 99STARTING - - 1000(7) - - - - -4O(min)TRANSIENT - 123% (7) 850 38,500(8) 102.6(8) 2200(7) - 0 to 104 (3)MAX REVERSE(9) - 750 - 88 1900 105 to 0 to 99

135

NOTES:(1) Torque limit applies within range of 1600-2000 propeller RPM (N 2). Below 1600 RPM, torque is limited to 49%.(2) Normal takeoff and maximum continuous operation oil pressure at gas generator speeds above 72% with oiltemperature between 60 and 71°C is 105 to 135 PSIG up to 21,000 feet. Above 21,000 feet, the minimum oil pressureis 85 PSIG. Plus or minus 10 PSIG fluctuations are acceptable. Oil pressure between 60 and 85 PSIG should betolerated only for the completion of the flight at power setting not to exceed 49% torque. Oil pressure below 60 PSIGare unsafe and require that either the engine be shut down or a landing be made as soon as possible, using theminimum power required to sustain flight. During extremely cold starts, oil pressure may reach 200 PSI.(3) These values are time limited to 5 minutes.(4) Cruise torque values vary with altitude and temperature.(5) At approximately 70% N1 .(6) High TGT at ground idle may be corrected by reducing accessory load and/or increasing N1 RPM.(7) These values are time limited to 5 seconds.(8) These values are time limited to 10 seconds.(9) This operation is time limited to 1 minute.(10) For every 5.6ºC below -48ºC ambient temperature, reduce maximum allowable N1 by 1.6%.

TABLE 5-2. GENERATOR LIMITS

GENERATOR LOAD

0 to 50%50 to 80%80 to 100%

*Right engine only

MINIMUM GAS GENERATOR RPM - N1WITHOUT AIR CONDITIONING *WITH AIR CONDITIONING

53% 60%60% 65%63% 70%

5-7

TM 55-1510-221-10

Section IV. LOADING LIMITS

5-17. CENTER OF GRAVITY LIMITATIONS.

Center of gravity limits and instructions forcomputation of the center of gravity are containedin Chapter 6. The center of gravity range will remainwithin limits, providing the aircraft loading isaccomplished according to instructions in Chapter 6.

The ability to sustain loss of engine powerand successfully stop, continue the take-off, or climb before or after gear retrac-tion is not assured for all conditions.Thorough mission planning must beaccomplished prior to takeoff by analysisof maximum takeoff weight permitted bytakeoff distance, accelerate-stop, positiveone-engine-inoperative climb at lift off,accelerate-go, takeoff climb gradient, andclimb performance. These data willdescribe performance capabilities for crit-ical mission decisions.

5-18. WEIGHT LIMITATIONS.

Maximum takeoff gross weight is 15,000pounds. Maximum landing weight is 15,000 pounds.Maximum ramp weight is 15,090 pounds. Maxi-mum zero fuel weight is 11,500 pounds.

Section V. AIRSPEED LIMITS, MAXIMUM AND MINIMUM

5-19. AIRSPEED LIMITATIONS. 5-23. WING FLAP EXTENSION SPEEDS.

Airspeed indicator readings contained in proce-dures, text, and illustrations throughout this Opera-tor’s Manual are given as indicated airspeed (IAS).Airspeed indicator markings (fig. 5-1) and placardedairspeeds, located on the cockpit overhead controlpanel (fig. 2- 12), are also indicated airspeeds.

The airspeed limit for APPROACH extension(40%) of the wing flaps is 198 KIAS. The airspeedlimit for full DOWN extension (100%) of the wingflaps is 153 KIAS. If wing flaps are extended abovethese speeds, the flaps or their operating mecha-nisms may be damaged.

5-20. MAXIMUM ALLOWABLE AIRSPEED.

The maximum allowable airspeed is 243 KIAS/0.42 Mach.

5-24. MINIMUM SINGLE-ENGINE CONTROL AIR-SPEED (Vmc ) .

5-21. LANDING GEAR EXTENSION SPEED.

The airspeed limit for extending the landing gearand for flight with the landing gear extended is 180KIAS.

Chapter 7, Section X describes minimum single-engine control airspeeds. The minimum single-engine control airspeed (Vmc ) at sea level standardconditions is 91 KIAS.

5-22. LANDING GEAR RETRACTION SPEED.5-25. MAXIMUM DESIGN MANEUVERINGSPEED.

The airspeed limit for retracting the landing gearis 162 KIAS.

The maximum design maneuvering speed is 168KIAS.

5-8

TM 55-1510-221-10

FLIGHT ENVELOPE CHART

Figure 5-2. Flight Envelope

Change 4 5-9

TM 55-1510-221-10

Section VI. MANUEVERING LIMITS

5-26. MANEUVERS.

For abrupt maneuvers above 168 KIAS. refer to FlightEnvelope (tip. 5-2).

a. The following maneuvers are prohibited.

1. Spins.

4. Any maneuver which results in a positive loadfactor of 3.06G's or a negative load factor of1.224G’s with wing flaps up, or a positive loadfactor of 2.0G’s or a negative 1224G’s withwing flaps down.

b. Recommended turbulent air penetration airspeed is158 KIAS.

3. Aerobatics of any kind. 5-27. BANK AND PITCH LIMITS.

3. Abrupt maneuvers above 168 KIAS. a. Bank limits are 60° left or right.

b. Pitch limits are 30° above or below the horizon.

Section VII. ENVIRONMENTAL RESTRICTIONS

5-28. ALTITUDE LIMITATIONS.

The maximum altitude that the aircraft may be operatedat is 31,000. When operating with inoperative yaw damp,the altitude limit is 17,000 feet.

5-29. TEMPERATURE LIMITS.

a. The aircraft shall not be operated when the ambienttemperatures are warmer than ISA 37°C at SL to 25,000feet. ISA 31°C above 25,000 feet.

b . Engine ice vanes shall be retracted at 1°C and above.

5-30. FLIGHT UNDER IMC (INSTRUMENTMETEOROLOGICAL CONDITIONS).

This aircraft is qualified for operation in instrumentmeteorological conditions.

5-30A. ICING LlMlTATlONS (TYPICAL).

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. Ice

accumulation 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 1/2-inch accumulation.

2. A 30 percent increase in torque per engine requiredto maintain a desired airspeed in level flight (not to exceed85 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.

5-10 Change 5

TM 55-1510-221-10

5-30B. ICING LIMITATIONS (SEVERE).

WARNING

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 in 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 certficated 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.

5-31. CROSSWIND LIMITATION.

The maximum crosswind component is 25 knots at 90°.The maximum angle of bank in a slip during landing is 8°.Landing the aircraft in a crab will impose side loads on thelanding gear and should he recorded on the DA Form 240813-1. Refer to Chapter 8 for crosswind landing technique.

5-32. OXYGEN REQUIREMENTS.

A minimum ten minute supply of supplemental oxygenshall be available during flight at or above an altitude of25,000 feet based on the highest total aircraft oxygen flowrates.

In addition to the supply required by the informationin the above paragraph. sufficient oxygen will be carriedfor each flight. assuming a decompression will occur at thealtitude or point of flight that is most critical from thestandpoint of oxygen need. and that after decompressionthe aircraft will descend. in accordance with the emergencyprocedures. to a flight altitude that Will allow successfultermination of the flight. Following the decompression, thecabin pressure altitude is considered to be the same as theflight altitude.

An oxygen system data/duration table may be found inChapter 2.

5-33. CABIN PRESSURE LIMITS.

Maximum cabin differential is 6.2 PSI.

5-34. CRACKED CABIN WINDOW / WIND-SHIELD.

If a crack occurs in an outer cabin window. the aircraftis limited to an altitude of 25,000 feet, and maximum cabinpressure differential is limited to 4.6 PSI. Maximumoperating time with a crack in an outer cabin window is 20hours. If an external windshield crack is noted. no action isrequired in flight. If an external crack occurs in eithercabin window or the windshield, refer to Chapter 9.Emergency Procedures.

Change 5 5-10.1/(5-10.2 blank)

TM 55-1510-221-10

Section VIII. OTHER LIMITATIONS

5-35. MAXIMUM DESIGN SINK RATE.

The maximum design sink rate is 500 feet perminute.

5-36. INSTRUMENT LANDING SYSTEM LIMITS.

During ILS approach do not operate the propel-lers in the 1750 to 1850 RPM range.

5-37. INTENTIONAL ENGINE OUT SPEED.

Inflight engine cuts below the safe one-engineinoperative speed (Vsse - 109 KIAS) are prohibited.

5-38. LANDING ON UNPREPARED RUNWAY.

Except in an emergency, propellers shouldbe moved out of reverse above 40 knotsto minimize propeller blade erosion, andduring crosswind to minimize stressimposed on propeller, engine and air-frame. Care must be exercised whenreversing on runways with loose sand ordust on the surface. Plying gravel willdamage propeller blades and dust mayimpair the pilot’s forward visibility at lowaircraft speeds.

5-39. MINIMUM OIL TEMPERATURE REQUIREDFOR FLIGHT.

Engine oil is used to heat the fuel upon enteringthe fuel control. Since no temperature measurementis available for the fuel at this point, it must be

assumed to be the same as the OAT. The minimumoil temperature graph (fig. 5-3) is provided for use asa guide in preflight planning, based on known orforecast operating conditions, to allow the operatorto become aware of operating temperatures whereicing at the fuel control could occur. If the plotshould indicate that oil temperatures versus OATare such that ice formation could occur during take-off or in flight, anti-icing additive per MIL-I-27686should be mixed with the fuel at refueling to insuresafe operation. In the event that authorized fuels(Prist) are not available the limitation of this chartapply.

Anti-icing additive must be properlyblended with the fuel to avoid deteriora-tion of the fuel cell. The additive concen-tration by volume shall be a minimum of0.060% and a maximum of 0.15%.Approved procedure for adding anti-icingconcentrate is contained in Chapter 2,Section XII.

JP-4 fuel per MIL-T-5624 has anti-icingadditive per MIL-I-27686 blended in thefuel at the refinery and no further treat-ment is necessary. Some fuel suppliersblend anti-icing additive in their storagetanks. Prior to refueling, check with thefuel supplier to determine if fuel has beenblended. To assure proper concentrationby volume of fuel on board, blend onlyenough additive for the unblended fuel.

5-11

TM 55-1510-221-10

MINIMUM OIL TEMPERATURE REQUIRED FOROPERATION WITHOUT ANTI-lClNG ADDlTlVE

0-60 -50 -40 -30 -20 -10 0

FUEL TEMPERATURE (OAT) ~ ºC

Figure 5-3. Minimum Oil Temperature

Section IX. REQUIRED EQUIPMENT FOR VARIOUS CONDITIONS OF FLIGHT

5-40. REQUIRED EQUIPMENT LISTING.

a. A Required Equipment for Various Condi-tions of Flight listing (fig. 5-4) is provided to enablethe pilot to indentify those systems/componentsrequired for flight. For the sake of brevity, the list-ing does not include obviously required items suchas wings, rudders, flaps, engines, landing gear, etc.Also the list does not include items which do notaffect the airworthiness of the aircraft such as galleyequipment, entertainment systems, passenger conve-nience items, etc. However, it is important to notethat ALL ITEMS WHICH ARE RELATED TOTHE AIRWORTHINESS OF THE AIRCRAFTAND NOT INCLUDED ON THE LIST AREAUTOMATICALLY REQUIRED TO BE OPERA-TIVE.

b. It is the final responsibility of the pilot todetermine whether the lack or inoperative status ofa piece of equipment on his aircraft will limit theconditions under which he may operate the aircraft.

(-) Indicates item may be inoperative for thespecified flight condition,

(*) Refers to remarks and/or exceptions columnfor explicit information or reference.

Numbered items contained in ( ) arerequired for flights by AR 95-l.

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.

5-12

TM 55-1510-221-10

SYSTEM and/or COMPONENT

ELECTRICAL POWER

1. AC Volts/Frequency Meter2. Battery3. Battery Charge Monitor System and Annuncia

tor4. DC Generator5. DC Generator Annunciator6. DC Load Meter

7. lnverter8. lnverter Annunciator

ENVIRONMENTAL

1. Bleed Air Fail Annunciators

2. Altitude Warning Annunciator (cabin)

3. Cabin Rate of Climb Indicator4. Differential Pressure/Cabin Altitude Indicator5. Duct Overtemp Annunciator6. Outflow Valve7. Pressurization Controller8. Safety Valve9. Bleed Air Shutoff Valve

FIRE PROTECTION

1. Engine Fire Detector System and Annunciator

FLIGHT CONTROLS

1. Flap Position Indicator

2. Flap System

3. Stall Warning Horn

4. Trim Tab Position Indicator(Rudder, Aileron, Elevator)

5. Yaw Damp System

111

222

21

2

1

1111112

2

1

1

1

3

1

111

222

21

2

1

1111112

2

1

1

1

3

1

111

222

21

2

1

1111112

2

1

1

1

3

1

111

222

21

2

1

11111

12

2

1

1

1

3

1

111

222

21

2

1

1111112

2

1

1

1

3

1

One may be inoperative providedcorresponding loadmeter is monitored.

May be inoperative provided bothinverters are operative.

Provided bleed air is not used fromside of failed light.May be inoperative provided air-plane remains unpressurized.

May be inoperative provided thatthe flap travel is visually inspectedprior to takeoff.

May be inoperative provided thatthe trim tabs are checked in theneutral position prior to each take offand checked for full range of opera-tion.May be inoperative for flight at andbelow 17,000 feet.

Figure 5-4. Required Equipment Listing (1 of 3)

5-13

TM 55-1510-221-10

Figure 54. Required Equipment Listing (2 of 3)

5-14

TM 55-1510-221-10

Figure 5-4. Required Equipment Listing (3 of 3)

5-15/(5-16 blank)

TM 55-1510-221-10

CHAPTER 6

WEIGHT/BALANCE AND LOADING

Section I. GENERAL.

6-1. EXTENT OF COVERAGE. of Class 1 aircraft forms and records are contained

Sufficient data has been provided so that, know-in DA PAM 738-751 and TM 55-1510-342-23.

ing the basic weight and moment of the aircraft, anycombination of weight and balance can be com-puted.

6-3. AIRCRAFT COMPARTMENT AND STA-TIONS.

6-2. CLASS.

Army Model RC-12H aircraft are in Class 1.Additional directives governing weight and balance

The aircraft is separated into two compartmentsassociated with loading. These compartments are thecockpit and the cabin. Figure 6-l shows the generaldescription of aircraft compartments.

Section II. WEIGHT AND BALANCE

6-4. PURPOSE.

The data to be inserted on weight and balancecharts and forms are applicable only to the individ-ual aircraft, the serial number of which appears onthe title page of the booklet entitled WEIGHT ANDBALANCE DATA supplied by the aircraft manufac-turer and on the various forms and charts whichremain with the aircraft. The charts and formsreferred to in this chapter may differ in nomencla-ture and arrangement from time to time, but theprinciple on which they are based will not change.

6-5. CHARTS AND FORMS.

The standard system of weight and balance con-trol requires the use of several different charts andforms. Within this Chapter, the following are used:

a. Chart C - Basic Weight and BalanceRecord, DD Form 365-3 (fig. 6-2).

b. Form F - Weight and Balance ClearanceForm F, DD Form 365-4 (Tactical), fig. 6-3).

6-6. RESPONSIBILITY.

The aircraft manufacturer inserts all aircraftidentifying data on the title page of the booklet enti-tled WEIGHT AND BALANCE DATA and on thevarious charts and forms. All charts, including onesample Weight and Balance Clearance Form F, ifapplicable, are completed at time of delivery. Thisrecord is the basic weight and balance data of the

aircraft at delivery. All subsequent changes in weightand balance are compiled by the weight and balancetechnician.

6-7. WEIGHT DEFINITIONS.

Weight definitions are as follows:

a. Basic Weight, The basic weight of an air-craft is that weight which includes all fixed operat-ing equipment and unusable fuel and engine oil. Itis only necessary to add variable or expendable loaditems for various missions. The basic weight of anaircraft varies with structural modifications andchanges in fixed operating equipment. The termbasic weight, when qualified with a word indicatingthe type of missions such as Basic Weight for Com-bat, Basic Weight for Ferry, etc., may be used inconjunction with directives stating what the equip-ment will be for these missions. For example, extrafuel tanks and various items of equipment installedfor long range ferry flight, which are not normallycarried on combat missions, will be included inBasic Weight for Ferry but not in Basic Weight forCombat.

b. Operating Weight. The operating weight isthe basic weight of the aircraft, including the crewand all equipment required for the mission, but notincluding fuel or payload.

c. Gross Weight. The gross weight is the totalweight of an aircraft contents, and fuel.

6-1

TM 55-1510-221-10

Figure 6-1. Aircraft Compartments and Stations

6-2

1.) The takeoff gross weight is the operat-ing weight plus the variable and expendable loaditems which vary with the mission.

(2.) The landing gross weight is the take-off gross weight minus the expended load items.

6-8. BALANCE DEFINITIONS.

Balance definitions are as follows:

a. Reference Datum. The reference datum isan imaginary vertical plane at, or forward of, thenose of the aircraft from which all horizontal dis-tances are measured for balance purposes. Diagramsof each aircraft show this reference datum as fuse-lage station zero.

b. Arm. Arm, for balance purposes, is the hor-izontal distance in inches from the reference datumto the center of gravity of the item. Arm may bedetermined from the Aircraft Compartment and Sta-tion Diagram (fig. 6-l).

c. Moment . Moment is the product of aweight multiplied by its arm. Moment divided by aconstant is generally used to simplify balance calcu-lations by reducing the number of digits. For thisaircraft, inches and moment/100 have been used.

d. Average Arm. Average arm is the armobtained by adding the weights and the moments ofa number of items and dividing the total moment bythe total weight.

e. Basic Moment. Basic moment is the sum ofthe moments of all items making up the basicweight. When using data from an actual weighing ofan aircraft, the basic moment is the total moment ofthe basic aircraft with respect to the referencedatum.

f Center of Gravity (CG). Center of gravity isthe point about which an aircraft would balance ifsuspended. Its distance from the reference datum isfound by dividing the total moment by the totalweight of the aircraft.

g. CG Limits. CG limits are the extremes ofmovement which the CG can have without makingthe aircraft unsafe to fly. The CG of the loaded air-craft must be within these limits at takeoff, in theair, and on landing.

6-9. CHART C - BASIC WEIGHT AND BALANCERECORD.

Chart C is a continuous history of the basicweight and moment resulting from structural andequipment changes made in service. At all times, the

TM 55-1510-221-10

last weight and moment/100 entry is considered thecurrent weight and balance status of the basic air-craft (fig. 6-2).

6-10. WEIGHT AND BALANCE CLEARANCEFORM F, DD FORM 365-4 (TACTICAL).

Form F (fig. 6-3) is a summary of the actual dis-position of load in the aircraft. It records the bal-ance status of the aircraft step by step. It serves asa work sheet to record weight and balance calcula-tions and any corrections that must be made toinsure that the aircraft will be within weight and CGlimits. It is necessary to complete a Form F prior toflight when an aircraft is loaded in a manner forwhich no previous valid Form F is available. A copymust remain in the aircraft for the duration of theflight. Form F (Tactical) is completed as follows:

1.

2.

3.

4.

Insert necessary identifying information att o p o f f o r m . I n b l a n k s p a c e s o f LIMITATIONS table, enter gross weightsfor takeoff and landing obtained from theWEIGHT LIMITATIONS paragraph inChapter 5.

Ref 1. Enter aircraft basic weight and indexor mom/100 figure. Obtain this informationfrom last entry on Chart C (fig. 6-2).

Ref 2. Leave blank (oil is included in basicweight).

Ref 3. Using compartment letterdesignations as s h o w n i n AircraftCompartment and Station Diagram (fig.6-1) enter number, weight, and mom/100figures of crew at their takeoff positions. Useactual crew weights if available. Enter totalof each compartment in WEIGHT andMOM/100 columns. To determine MOM/100 of crew use Table 6-1, Occupants UsefulLoad, Weights and Moments.

NOTE

The maximum baggage compartmentweight is 410 pounds. Also the floor load-ing limit of 100 Lbs/Sq Ft shall not beexceeded.

5. Ref 4. Enter sum of weights and sum ofmom/100 figures for Ref 1 through Ref 3 toobtain OPERATING WEIGHT and corre-sponding mom/ 100 figure.

6. Ref 5. Enter the item description (flare/chaff), total amount, weight, and MOM/100of all expendable stores.

6-3

TM 55-1510-221-10

Figure 6-2. Chart C - Basic Weight and Balance Record, DD Form 365-3

7. Ref 6. Not applicable to RC-12H aircraft.

8. Ref 7. Enter the number of gallons, weightand sum of mom/100 figures for takeofffuel. The weight of fuel used in warm up andtaxiing should not be included. If JP-4 fuelat a density of 6.5 Lb/Gal is being used, useFuel Moments Table (Table 6-2) to deter-mine fuel weight and mom/ 100 figures. Ifother than JP-4 is being used, use the follow-ing procedures to obtain accurate fuel weightand mom/100 figures from the Fuel MomentChart and to record information on Form F.

a. Assume 6.7 Lb/Gal for JP-5 fuel or 6.0Lb/Gal for AVGAS.

b. Multiply the total number of gallons offuel in the aircraft times the computedfue l dens i t y f i gu re and t he r ebydetermine actual fuel weight.

c. Use Fuel Moment Table (Table 6-2) todetermine mom/ 100 figure.

d. Enter the weight and correspondingmom/100 figure in the correspondingcolumns of Ref 7. Also, enter a figurefor the total fuel gallons known to be inthe aircraft.

9. Ref 8. Not applicable to RC-12H aircraft.

10. Ref 9. Enter sum of weights and momentsfor reference 4 through reference 8 oppositeTAKEOFF CONDITION (uncorrected).

11. Ref 10. Enter TAKEOFF C.G. (uncorrected)as determined from weight and momentvalues of reference 9.

Check WEIGHT figure opposite Ref 10 againstGROSS WT. TAKEOFF. Check mom/100 figureopposite Ref 10 with Center-of-Gravity LimitsTable (Table 6-4) to ascertain that CG is withinallowable limits.

12. Ref 11. If changes in amount or distributionof loads are required, indicate necessaryadjustments by proper entries in CORREC-TIONS table. Enter a brief description ofadjustment made in column marked ITEM.Add all weight and moment decreases andinsert totals in space opposite TOTALWEIGHT REMOVED. Add all weight andmoment increase and insert totals in spaceopposite TOTAL WEIGHT ADDED. Sub-tract smaller from larger of two totals andenter differences (with applicable + or -sign) opposite NET DIFFERENCE. Transferthese NET DIFFERENCE figures to spacesopposite Ref 11.

6-4

TM 55-1510-221-10

Figure 6-3. Weight and Balance Clearance, DD Form 365-4 (Tactical)

6-5

TM 55-1510-221-10

Table 6-1. Occupants Useful Load, Weights andMoments

USEFUL LOAD WEIGHTSAND MOMENTS

OCCUPANTS

CREWWEIGHT F.S.129

MOM/100

80 10390 116100 129110 142120 155130 168140 181150 194160 206170 219180 232190 245200 258210 271220 284230 297240 310250 323

13. Ref 12. Enter sum of the difference betweenRef 10 and Ref 11. Recheck against GROSSWT. TAKEOFF in LIMITATIONS table toassure that this figure does not exceedallowable limit.

Record figures in LIMITATIONS table as fol-lows: In FORWARD and AFT space for PERMISSI-BLE C.G. TAKEOFF enter an inches figureobtained as follows: Match weight figure recorded inRef 12 with a corresponding figure in the GROSSWEIGHT-POUNDS on Center-of-Gravity LimitsTable (Table 6-4). Determine the FORWARD limitin inches figure for the weight matched and recordin the FORWARD space stated. Enter AFT C.G.LIMIT in inches in AFT space.

14. Ref 13. By referring to Center-of-GravityLimits Table (Table 6-4), determine takeoffC.G. position. Enter this figure in space pro-vided opposite TAKEOFF C.G. in inches.Insure that this position is within the FWDand AFT C.G. limit in LIMITATIONStable.

15. Ref 14. Less expendables.

Fuel.

a. Enter estimated weight of fuel to beexpended. Subtract this figure fromweight of fuel on board (reference 7).This figure represents the landing fuelweight. Use the Fuel Moments Table(Table 6-2) to determine landing fuelmoment. Subtract the landing fuelmoment from the total fuel moment onboard (reference 7). This figure repre-sents the moment of fuel expended.Enter in reference 14.

NOTE

Do not consider reserve fuel as expendedwhen determining ESTIMATED LAND-ING CONDITIONS.

Flare/Chaff cartridges.

b. Determine total weight of flares and/orchaff cartridges t h a t h a v e b e e nexpended. Use the Aircraft Surviva-bility Equipment table (Table 6-3) tode t e rmine l and ing expendab l e smoment. Enter figures in WEIGHTand MOM/ 100 columns.

c. Ref 15. Enter differences in weightsand MOM/100 figures between refer-ence 12 and totals of reference 14.

Record figures in LIMITATIONS table as fol-lows: In FORWARD and AFT space for PERMISSI-BLE C.G. LANDING, enter a figure obtained as fol-lows: Match weight figure recorded in reference 15with a corresponding figure in the GROSSWEIGHT-POUNDS on Center-of-Gravity LimitsTable (Table 6-4). Determine the FORWARD limitin inches figure for the weight matched and recordin the FORWARD space stated. Within the AFTspace for PERMISSIBLE C.G. LANDING, recordthe AFT C.G. limit in inches. Check data againstPERMISSIBLE C.G. LANDING in LIMITATIONStable.

d. Ref 16. Refer to Center-of-GravityLimits Table (Table 6-4) to determinelanding C.G. position. Enter this figurein space provided opposite ESTI-MATED LANDING C.G. in inches.

Check Ref 16 against PERMISSIBLE C.G.LANDING in LIMITATIONS Table.

Necessary signatures must appear at bottom ofform.

6-6

TM 55-1510-221-10

F U E L D E N S I T Y / W E I G H T

v s

T E M P E R A T U R EMODEL: UC-12BDATE: 14 MAY 1979DATA BASIS: FLIGHT TESTCONFIGURATION:

ENGINE: PT6A-41PROPELLER: T10178FUEL GRADE: JP-5FUEL DENSITY: 6.8 LB/GALEXAMPLE:FUEL

TEMPERATURE: . . . . . . . . . . . . . . . . 28°CFUEL GRADE: . . . . . . . . . . . . . . . . . . . . . JP-5SPECIFIC WEIGHT . . . . . . . . . . . . . . . . = 6.7 LB/US GALFUEL QUANTITY: . . . . . . . . . . . . . . . . . . 130 US GALFUELWEIGHT: . . . . . . . . . . . . . . . . . . . (6.7 X 130) = 871 LBS

AP 004484

Figure 6-4. Density Variation of Aviation Fuel.

6-7

TM 55-1510-221-10

Table 6-2. Fuel Center-of-Gravity Moments

USEFUL LOAD WEIGHTS AND MOMENTSUSABLE FUEL

6-8

TM 55-1510-221-10

Table 6-3. Survivability Equipment Weights and Moments

ItemNACELLE DISPENSER

Weight(lb)

Moment/100

Dispenser (Empty)Chaff Cartridges (30)

TOTAL : (Dispenser/30 chaff cartridges)

FUSELAGE DISPENSER

10 219 19

I19

I40

Dispenser (Empty) 10 49Chaff and/or Flare Cartridges 9 2 6

TOTAL : (Dispenser/30 chaff and/or flare 19 55cartridges)

Change 4 6-9

TM 55- 1510-221-10

Table 6-4. Center-of-Gravity Moments (sheet 1 of 3)

CENTER OF GRAVITY MOMENT TABLE - MOMENT/100

6-10

TM 55-1510-221-10


CENTER OF GRAVITY MOMENT TABLE - MOMENT 100

6-11

TM 55-1510-221-10


CENTER OF GRAVITY MOMENT TABLE - MOMENT/ 100 (CONT’D)

6 - 1 2

TM 55-1510-221-10

Table 6-5. C.G. Limits (Landing Gear Down) - Restricted Category

*CENTER OF GRAVITY LIMITS (LANDING GEAR DOWN) RESTRICTED CATEGORY

NOTES:-The moment/100 for retraction of the alighting gear is - 60.4. Loadings based on wheels-down condition which fall withinthe limiting moments in the table, will be satisfactory for flight with alighting gear retracted.

6-13

TM 55-1510-221-10

Section III. FUEL/OIL

6-11. FUEL LOAD. 6-12. FUEL AND OIL DATA.

Fuel loading imposes a restriction on theamount of load which can be carried. The requiredfuel must first be determined, then that weight sub-tracted from the total weight of crew and fuel.Weight up to and including the remaining allowablecapacity can be subtracted directly from the weightof crew and fuel. As the fuel load is increased, theloading capacity is reduced.

a. Fuel Moment Table. This table (Table 6-2)shows fuel moment/ 100 given US gallons or poundsfor JP-4 and JP-5.

b. Oil Data. Total oil weight is 62 pounds andis included in the basic weight of the aircraft. Servic-ing information is provided in Section XII of Chap-ter 2.

Section IV. CENTER OF GRAVITY

6-13. CENTER OF GRAVITY LIMITATIONS.

Removal of mission gear may result inexceeding the forward center-of-gravitylimit.

Center of gravity limitations are expressed inARM inches which refers to a positive measurementfrom the aircraft’s reference datum. The forward CGlimit at 11,279 Lbs. or less is 181.0 ARM inches.The forward-sloping CB limit line from 11,279 Lbs.to 13,500 Lbs., and straight up to 15,000 Lbs., isfuselage station 188.3. At 15,000 Lbs. or less, the aftCG limit is 195.1 ARM inches. The Center of Grav-ity Limitations Table (Table 6-4) is designed toestablish forward and aft CG limitations.

Section IV. Cargo Loading

6-14. LOAD PLANNING. The basic factors to beconsidered in any loading situation are as follows:

a. Cargo shall be arranged to permit access toall emergency equipment and exits during flight.

b. Floorboard structural capacity shall be con-sidered in the loading of heavy or sharp-edged con-tainers and equipment. Shorings shall be used to dis-tribute highly condensed weights evenly over thecargo areas.

c. All cargo shall be adequately secured toprevent damage to the aircraft, other cargo, or theitem itself.

6-15. LOADING PROCEDURE.

NOTE

The cabin door is weight limited to amaximum of 300 pounds to prevent pos-sible structural damage.

Loading of cargo is accomplished through thecabin door (21.5 in. X 50.0 in.) or the cargo door(52.0 in. X 52.0 in).

6-16. SECURING LOADS.

All cargo shall be secured with restraints strongenough to withstand the maximum force exerted inany direction. The maximum force can be deter-mined by multiplying the weight of the cargo itemby the applicable load factor. These established loadfactors (the ratio between the total force and theweight of the cargo item) are 1.5 to the side andrear, 3.0 up, 6.6 down, and 9.0 forward.

6-14

TM 55-1510-221-10

C H A P T E R 7

P E R F O R M A N C E

TABLE OF CONTENTS

Introduction to Performance .......................................................................................................................... 7-lTakeoff Flight Path .......................................................................................................................................... 7-3Comments Pertinent to the use of Performance Graphs .............................................................................. 7-6Table of Contents ............................................................................................................................................ 7-8Airspeed Calibration - Normal System ........................................................................................................ 7-10Altimeter Correction - Normal System ........................................................................................................ 7-11Airspeed Calibration - Alternate System ...................................................................................................... 7-12Altimeter Correction - Alternate System ...................................................................................................... 7-13Free Air Temperature Correction ................................................................................................................ 7-14ISA Conversion .............................................................................................................................................. 7-15Fahrenheit to Celsius Temperature Conversion .......................................................................................... 7- 16Minimum Takeoff Power at 2000 RPM ...................................................................................................... 7-17Maximum Takeoff Weight Permitted by Enroute Climb Requirement ...................................................... 7-18Takeoff Weight to Achieve Positive One-Engine-Inoperative Climbat Lift-Off - Flaps 0% .................................................................................................................................... 7-19Takeoff Weight to Achieve Positive One-Engine-Inoperative Climbat Lift-Off - Flaps 40% .................................................................................................................................. 7-20Wind Components ........................................................................................................................................ 7-21Takeoff Distance - Flaps 0% ........................................................................................................................ 7-22Accelerate-Stop - Flaps 0% ............................................................................................................................ 7-23Accelerate-Go Distance Over 50 FT Obstacle - Flaps 0% .......................................................................... 7-24Takeoff Climb Gradient - One-Engine-Inoperative - Flaps 0% .................................................................. 7-25Takeoff Distance - Flaps 40% ...................................................................................................................... 7-26Accelerate-Stop - Flaps 40% .......................................................................................................................... 7-27Accelerate-Go Distance Over 50 FT Obstacle - Flaps 40% ........................................................................ 7-28Takeoff Climb gradient - One Engine Inoperative -Flaps 40% ...................................................................................................................................................... 7-29Climb - Two-Engine - Flaps 0% .................................................................................................................. 7-30Climb - Two-Engine - Flaps 40% ................................................................................................................ 7-31Climb - One Engine Inoperative .................................................................................................................. 7-32Service Ceiling One-engine-Inoperative ........................................................................................................ 7-33Time, Fuel, and Distance to Cruise Climb .................................................................................................. 7-34Maximum Cruise Power 1900 RPM ISA-30°C .......................................................................................... 7-35Maximum Cruise Power 1900 RPM ISA-20°C .......................................................................................... 7-36Maximum Cruise Power 1900 RPM ISA-1OºC .......................................................................................... 7-37Maximum Cruise Power 1900 RPM ISA .................................................................................................. 7-38Maximum Cruise Power 1900 RPM ISA+10ºC ........................................................................................ 7-39Maximum Cruise Power 1900 RPM ISA+20ºC ........................................................................................ 7-40Maximum Cruise Power 1900 RPM ISA+30ºC ........................................................................................ 7-41Maximum Cruise Power 1900 RPM ISA+37ºC ........................................................................................ 7-42Maximum Cruise Speeds 1900 RPM .......................................................................................................... 7-43Maximum Cruise Power 1900 RPM ........................................................................................................ 7-44Fuel Flow At Maximum Cruise Power 1900 RPM .................................................................................. 7-45Range Profile - Maximum Cruise Power 1900 RPM ................................................................................ 7-46Maximum Range Power 1700 RPM ISA-30°C .......................................................................................... 7-47Maximum Range Power 1700 RPM ISA-20°C .......................................................................................... 7-48Maximum Range Power 1700 RPM ISA-10ºC .......................................................................................... 7-49

7-1

TM 55-1510-221-10

Maximum Endurance Power 1700 RPM ISA+20ºCMaximum Endurance Power 1700 RPM ISA+30ºC

................................................................................

Maximum Endurance Power 1700 RPM ISA+37ºC................................................................................

Range Profile - Long Range Power 1700 RPM................................................................................

Range Profile - 542 Gallons Usable Fuel........................................................................................

....................................................................................................Endurance Profile - 542 Gallons Usable FuelTime, Fuel, And Distance to Descend

............................................................................................

Climb - Balked Landing........................................................................................................

................................................................................................................................Normal Landing Distance Without Propeller Reversing- Flaps 100% ..................................................................................................................................................Landing Distance Without Propeller Reversing - Flaps 0% ......................................................................MaximumMaximumMaximumMaximumMaximumMaximumMaximumMaximumMaximumMaximum

Range Power 1700 RPM ISA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Range Power 1700 RPM ISA+ 10°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Range Power 1700 RPM ISA+20ºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Range Power 1700 RPM ISA+30ºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Range Power 1700 RPM ISA+37ºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Endurance Power 1700 RPM ISA-30°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Endurance Power 1700 RPM ISA-20°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Endurance Power 1700 RPM ISA-1OºC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Endurance Power 1700 RPM ISA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Endurance Power 1700 RPM ISA+ 10°C .,.........................................................................,....

7-607-617-62 7-637-647-657-667-67

7-687-697-507-517-527-537-547-557-567-577-587-59

7-2

TM 55-1510-221-10

C H A P T E R 7

P E R F O R M A N C E

7-l. INTRODUCTION TO PERFORMANCE.

The graphs in this Section present performanceinformation for takeoff, climb, cruise, and landing atvarious parameters of weight, altitude, and tempera-ture.

The following example presents calculations fora proposed 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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5333 feet1Altimeter Setting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30.02 in. HgWind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 030º at 13 knotsRunway 35R Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,000 feet1Route of trip:DEN - J116 - EKR - J173 - SLC - J154 -BAM - J32 - RN0Cruise Altitude:26,000 feetAt Cannon International (RNO):Free Air Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . 32°C (90°F)Field Elevation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4412 feet1Altimeter Setting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29.60 in. HgWind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200º at 15 knotsRunway 25 Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6101 feet1

1 Source: NOAA Standard Instrument Depar-tures for Western United States, 9 APR1987.

2 Source: NOAA Enroute High Altitude -U.S.Chart H-l, 9 APR 1987.

3 MEA on NOAA Enroute Low Altitude -U.S.Chart L-8, 9 JUN 1983.

4 Includes distance between airport andVORTAC, per NOAA Airport/FacilityDirectory (Southwest US.), 9 APR 1987.

7-3. PRESSURE ALTITUDE.

To determine the approximate pressure altitudeat origin and destination airports, add 1000 feet tofield elevation 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 set-ting FROM 29.92 in. Hg, then multiply the answerby 1000 to find the difference in feet between fieldelevation and pressure altitude.

Pressure Altitude at DEN:

29.92 in. Hg - 30.02 in. Hg = -0.10-0.10 x 1000 feet = -100 feetThe pressure altitude at DEN is 100 feetbelow field elevation.Pressure altitude at DEN = 5333 - 100 = 5233 feet.Pressure altitude at RNO:29.92 in. Hg - 29.60 in. Hg = 0.320.32 x 1000 feet = 320 feetThe pressure altitude at RN0 is 320 feetabove field elevation.Pressure altitude at RN0 = 4412 + 320 = 4732 feet.

7-4. PERFORMANCE EXAMPLE.

Maximum takeoff weight (from LIMITATIONSSection) = 15,000 pounds

7-5. MAXIMUM TAKEOFF WEIGHT PERMITTEDBY ENROUTE CLIMB REQUIREMENT.

Enter the graph at 5233 feet take-off field pres-sure altitude to 28°C takeoff FAT:

Maximum Allowable Takeoff Weight 14,200 pounds

The maximum takeoff weight permitted by theEnroute Climb Requirement graph is the only oper-ating limitation required to meet applicable FARrequirements. Information has been presented, how-ever, to determine the takeoff weight, field require-ments, and takeoff flight path assuming an enginefailure occurs during the take-off procedure. The fol-lowing illustrates the use of these charts.

7-6. TAKEOFF WEIGHT TO ACHIEVE POSITIVEONE-ENGINE-INOPERATIVE CLIMB AT LIFTOFF(Flaps 0%).

Enter the graph at 5233 feet to 28ºC, to deter-mine the maximum weight at which the acceler-ate-go procedure should be attempted.

Maximum Accelerate-Go Weight . . . . 13,480 pounds

7 - 3

TM 55-1510-221-10

7-7. ACCELERATE-STOP (FLAPS 0%).

Enter the Accelerate-Stop graph at 28ºC, 5233feet pressure altitude, 13,480 pounds, and 10 knotshead wind component:

Accelerate-Stop Distance . . . . . . . . . . . . . . . . . . . . . . . . 5050 feetTakeoff Decision Speed (VR

) . . . . . . . . . . . . . . . . . . . . 98 knots

7-8. TAKEOFF DISTANCE (FLAPS 0%).

Enter the graph at 28ºC, 5233 feet pressure alti-tude, 13,480 pounds, and IO knots head wind com-ponent:

Ground Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3420 feetTotal Distance Over 50-foot Obstacle . . . . 5100 feetTakeoff Speed:At Rotation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 knotsAt 50 Feet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 16 knots

7-9. TAKEOFF FLIGHT PATH EXAMPLE.

The following example assumes the aircraft isloaded so that takeoff weight is 10,000 pounds.

7-10. ACCELERATE-GO DISTANCE OVER 50-FOOT OBSTACLE (FLAPS 0%).

Enter the graph at 28°C, 5233 feet pressure alti-tude, 10,000 pounds, and 10 knots head wind com-ponent:

Total Distance Over 50-foot Obstacle . . . . 5550 feetSpeed at Rotation (VR) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 93 knotsSpeed at 35 Feet Above Runway (Climb Speed) Knots

7-11. TAKEOFF CLIMB GRADIENT - ONEENGINE INOPERATIVE (FLAPS 0%).

Enter the graph at 28°C, 5233 feet pressure alti-tude, and 10,000 pounds:

Climb Gradient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1%Climb Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 knots

A 5.1% climb gradient is 51 feet of verticalheight per 1000 feet of horizontal distance.

NOTE

The graphs for take-off climb gradientassume a zero-wind condition. Climbinginto a head wind will result in higherangles of climb, and hence better obstacleclearance capabilities.

Calculations of the horizontal distance to clearan obstacle 100 feet above the runway surface:

7-4

Horizontal distance used to climb from 50 feet to 100 feet -(100 - 50) (1000 ÷ 51) = 981 feetTotal Distance = 5500 + 981- 6531 feetResults are illustrated below:

7-12. FLIGHT PLANNING.

Calculations for flight time, block speed, andfuel requirements for a proposed flight are detailedbelow using the same conditions presented on page4-3, and a takeoff weight of 12,000 pounds.

Enter ISA CONVERSION graph at the condi-tions indicated:

DEN-BVL

Pressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26,000 feetFAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -20°CISA Condition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ISA + 17°C

BLV-RN0

Pressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26,000 feetFAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -10°CIAS Condition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ISA + 27°C

Enter the TIME, FUEL, AND DISTANCE TOCLIMB Graph at 28°C to 5233 feet, and to 12,000pounds, and enter at -10°C to 26,000 feet, and to 12,000 pounds, and read:

Time to Climb . . . . . . . . . . 30.0-3.1 = 26.9=27 minutesFuel Used to Climb . . . . . . . . . . . . . . . . 379.3-60.0=319.3=

320 poundsDistance Traveled . . . . . . 85.7-8.7=77 nautical miles

Enter the TIME, FUEL, AND DISTANCE TODESCEND Graph at 26,000 feet, and enter again at4732 feet, and read:

Time to Descend . . . . . . 17.7-3.1= 14.6 = 15 minutesFuel Used to Descend . . . . 198.5-50.0=149 poundsDistance Traveled . . . . . . . . . . . . . . . . . . . . . . . . 86.5- 15.3 =71.2=

71 nautical miles

An estimated average cruise weight of 11,200pounds was used for this example.

Enter the tables for MAXIMUM ENDURANCEPOWER 1700 RPM for ISA + 10°C ISA + 20°C,and ISA + 30°C and read the cruise speeds for 26,000 feet at 12,000 pounds and 11,000 pounds:

Cruise True Airspeed (ISA + 17°C) . . . . . . 169 knotsCruise True Airspeed (ISA + 27°C) . . . . . . 172 knots

Enter the *MAXIMUM ENDURANCEPOWER 1700 RPM Tables for ISA + 10°C ISA +2OºC, and ISA + 30°C at 12,000 pounds and 11,000pounds and interpolate the recommended torquesettings for ISA + 17°C and ISA + 27°C at 11,200pounds.

ISA + 17°C . . . . . . . . . . . . . . . . . . . . . . . . 40% torque per engine

TM 55-1510-221-10

Figure 7-1. Takeoff Flight Path

ISA + 27°C . . . . . . . . . . . . . . . . . . . . . . . . 41% torque per engine

Enter the *MAXIMUM ENDURANCEPOWER 1700 RPM Tables for ISA + 10°C ISA +20°C and ISA + 30°C at 12,000 pounds and 11,000pounds at 26,000 feet, and interpolate the fuel flowsfor ISA + 17°C and ISA + 27°C at 11,200 pounds.

ISA + 17°CFuel Flow Per Engine . . . . . . . . . . . . . . . . . . . . . . 198.75 Lbs/hrTotal Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 397.50 Lbs/hrISA + 27°CFuel Flow Per Engine . . . . . . . . . . . . . . . . . . . . . . . . 203.5 Lbs/hrTotal Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 407 Lbs/hr

NOTE

For flight planning, enter these charts atthe forcasted ISA condition; for enroutepower settings and fuel flows, enter at theactual indicated FAT.

Time and fuel used were calculated at MAXI-MUM ENDURANCE POWER 1700 RPM as fol-lows:

Time =Fuel Used =

Distance Ground SpeedDistance x Total Fuel Flow

Ground Speed

Results are as follows:

7-13. RESERVE FUEL.

Reserve Fuel is calculated as 45 minutes at Max-imum Range Power 1700 RPM. Use planned cruisealtitude (26,000 feet), forecasted ISA condition (ISA+ 27°C) and estimated weight at end of plannedtrip (10,309 pounds). (Since the lowest weight col-umn in the tables is 11,000 pounds, assume weightat the end of the planned trip to be 11,000 pounds,and use that fuel flow value for this example.)

Enter the tables for MAXIMUM RANGEPOWER 1700 RPM for ISA + 20°C and ISA +30°C at 11,000 Lbs and 26,000 feet, and read thetotal fuel flows:

ISA + 20°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 478 Lbs/hrISA + 30°C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 448 Lbs/hr

Then interpolate for the fuel flow at ISA + 27°Cas follows:

Change in Fuel Flow = 478 - 448 = 30 Lbs/hr

Change in Temperature = (ISA + 20°C) - (ISA+ 30°C) = 10°C

Rate of Change in Fuel Flow = Change in FuelFlow ÷ Change in Temperature

7-5

TM 55-1510-221-10

Rate of Change in Fuel Flow = (30 Lbs/hr) ÷(10°C)

Rate of Change in Fuel Flow = 3.0 Lbs/hrdecrease per 1°C increase

Temperature increase from ISA + 20°C to ISA+ 27°C = 7°C

Total Change in Fuel Flow = 7 x 3.0 Lbs/hr =21.0 Lbs/hr

Total Fuel Flow = (ISA + 20°C Fuel Flow) +(Total Change in Fuel Flow)

Total Fuel Flow = (478) - (21) = 457 Lbs/hr

Reserve Fuel = 45 minutes x Total Fuel Flow

Reserve Fuel = (0.75) x (457 Lbs/hr) = 342.75= 343 lbs.

Total Fuel Requirement = 1781 + 343 = 2124pounds

7-14. ZERO FUEL WEIGHT LIMITATION.

For this example, the following conditions wereassumed:

Ramp Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,090 poundsWeight of Usable Fuel Onboard . . . . . . 2124 pounds

Zero Fuel Weight = Ramp Weight - Weight ofUsable Fuel Onboard

Zero Fuel Weight = (12,090) - (2124) = 9966pounds

Maximum zero fuel weight limitation (fromLIMITATIONS section) = 11,500 pounds.

Maximum Zero Fuel Weight Limitation has notbeen exceeded.

Anytime the Zero Fuel Weight exceeds the Max-imum Zero Fuel Weight Limit, the excess must beoff-loaded from PAYLOAD. If desired, additionalFUEL ONLY may then be added until the rampweight equals the Maximum Ramp Weight Limit of15.090 Lbs.

7-15. LANDING INFORMATION.

The estimated Landing Weight is determined bysubtracting the fuel required for the trip from theRamp Weight:

Ramp Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,090 LbsFuel Required for Total Trip . . . . . . . . . . 1781 poundsLanding Weight (12,090 - 1781) . . . . 10,309 pounds

Enter the NORMAL LANDING DISTANCEWITHOUT PROPELLER REVERSING - FLAPS100% Graph at 32°C, 4732 feet, 10,309 pounds, and10 knots head wind component:

Ground Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1800 feetTotal Distance Over 50-foot Obstacle . . . . 2510 feetApproach Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99 knots

Enter the CLIMB - BALKED LANDING Graphat 32°C, 4732 feet, and 10,309 pounds:

Rate of Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1270 ft/minClimb Gradient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.2%

7-16. COMMENTS PERTINENT TO THE USE OFPERFORMANCE GRAPHS.

a . In addition to presenting the answer for aparticular set of conditions the example on a graphalso presents the order in ‘which the various scaleson the graph should be used. For instance, if thefirst item in the example is FAT, then enter thegraph at the existing FAT.

b. The reference lines indicate where to beginfollowing the guidelines. Always project to the refer-ence line first, then follow the guidelines to the nextknown item by maintaining the same PROPOR-TIONAL DISTANCE between the guide line aboveand the guide line below the projected line. Forinstance, if the projected line intersects the referenceline in the ratio of 30% down/70% up between theguidelines, then maintain this same 30%/70% rela-tionship between the guide lines and follow them tothe answer or next known item.

c. The associated conditions define the spe-cific 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. Indicated airspeeds (IAS) were obtainedusing the Airspeed Calibration - Normal Systemgraph.

f: Notes have been provided on variousgraphs and tables to approximate performance withice vanes extended. The effect will vary, dependingupon airspeed, temperature, altitude, and ambientconditions. At lower altitudes, where operation onthe torque limit is possible, the effect of ice vaneextension will be less, depending upon how muchpower can be recovered after the ice vanes havebeen extended.

7-6

AIRSPEED CALIBRATION - NORMAL

TM 55-1510-221-10

SYSTEM

Figure 7-2. Airspeed Correction - Normal System

7-7

TM 55-1510-221-10

ALTIMETER CORRECTION - NORMAL SYSTEM

Figure 7-3. Altimeter Correction - Normal System

7-8

TM

55-1510-221-10

Figure 7-4. A

irspeed Calibration - A

lternate System

7-9

TM

55-1510-221-10

Figure 7-5. A

ltimeter C

orrection - Alternate System

7-10

TM 55-1510-221-10

INDICATED OUTSIDE AIR TEMPERATURE CORRECTION

Figure 7-6. Free Air Temperature Correction

7-11

TM 55-1510-221-10

ISA CONVERSION

PRESSURE ALTITUDE vs FREE AIR TEMPERATURE

Figure 7-7. ISA Conversion

7-12

TM 55-1510-221-10

Figure 7-8,

Fahrenheit to

Celsius

Tem

perature C

onversion

7-13

TM 55-1510-221-10

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 65 KNOTS ATWHICH TAKEOFF PRESENTED IN THIS SECTION CAN BE REALIZED. ANY EXCESS POWER WHICHCAN BE DEVELOPED WITHOUT EXCEEDING ENGINE LIMITATIONS SHOULD BE UTILIZED.

3. FOR OPERATION WITH ICE VANES EXTENDED, INCREASE FIELD PRESSURE ALTITUDE 1000 FEETBEFORE ENTERING GRAPH.

Figure 7-9. Minimum Takeoff Power at 2000 RPM

7-14

TM 55-1510-221-10

MAXIMUM TAKEOFF WEIGHT PERMITTED BY ENROUTE

ASSOCIATED CONDITIONS:CLIMB REQUIREMENT

EXAMPLE:POWER ....................... MAXIMUM CONTINUOUSINOPERATIVE PROPELLERS ........ FEATHEREDLANDING GEAR ...................... UPFLAPS . . . . . . . . . . . . . . . . . . . . . . . 0% MAXIMUM TAKEOFF

WEIGHTNOTE: ONE-ENGINE-INOPERATIVE PERFORMANCE WEIGHT LIMIT IS FOR

RATE-OF-CLIMB CAPABILITIES AT 5000 FT PRESSURE ALTITUDE.REFER TO THE l CLIMB - ONE ENGINE INOPERATIVE’ GRAPHFOR ACTUAL CLIMB CAPABILITIES APPLICABLE TO THEPARTICULAR TEMPERATURE AND ALTITUDE BEING CONSIDERED.

PRESSURE ALTITUDE . . . . . . 5233 FTFAT . . . . . . 28°C

14.200 LBS

Figure 7-10. Maximum Takeoff Weight Permitted by Enroute Climb Requirement

7-15

TM 55-1510-221-10

TAKEOFF WEIGHT - FLAPS 0%TO ACHIEVE POSITIVE ONE-ENGINE-INOPERATIVE CLIMB AT LIFT-OFF

ASSOCIATED CONDITIONS:

POWER . . . TAKEOFFFLAPS 0%

EXAMPLE:

PRESSURE ALTITUDEFAT

. . . . 5233 FT

...... 28°CLANDING GEAR .... . . . . . . . . . . . . DOWNINOPERATIVE PROPELLER FEATHERED TAKEOFF WEIGHT ........ . . . 13,480 LBS....

Figure 7-11. Takeoff Weight - Flaps 0%, To Achieve Positive One-Engine Climb at Lift-Off

7-16

TM 55-1510-221-10

TAKEOFF WEIGHT - FLAPS 40%TO ACHIEVE POSITIVE ONE-ENGINE-INOPERATIVE CLIMB AT LIFT-OFF

Figure 7-12. Takeoff Weight - Flaps 40%, To Achieve Positive One-Engine Climb at Lift-Off Flaps 40%

7-17

TM 55-1510-221-10

WIND COMPONENTSDemonstrated Crosswind is 17 kts

EXAMPLE:

WIND SPEEDANGLE BETWEEN WIND DIRECTION AND FLIGHT PATHHEADWIND COMPONENTCROSSWIND COMPONENT

Figure 7-13. Wind Components

7-18 Change 4

20 KTS50%

13 KTS15 KTS

TM 55-1510-221-10

Figure 7-14.Takeoff D

istance - Flaps 0%

Change 4

7-19

TM

55-1510-221-10

Figure 7-15. Accelerate-Stop - Flaps 0%

7-20C

hange 4

TM

55-1510-221-10

Figure 7-16. Accelerate-G

o Distance O

ver 50-FT O

bstacle - Flaps 0%

7-21

TM 55-1510-221-10Figure 7-17. T

akeoff Clim

b Gradient - O

ne-Engine-Inoperative - Flaps 0%

7-2

2

TM 55-1510-221-10

Figure 7-18. Takeoff D

istance - Flaps 40%

7-23

TM 55-1510-221-10

7-24

Figure 7-19. Accelerate-Stop - Flaps 40%

TM 55-1510-221-10

Figure 7-20. Accelerate-G

o Distance O

ver 50-FT O

bstacle - Flaps 40%

7-25

TM 55-1510-221-10

Figure 7-21. Takeoff C

limb gradient - O

ne-Engine-Inoperative - Flaps 40%

7-26

TM 55-1510-221-10

Figure 7-22. Clim

b Tw

o-Engine - Flaps 0%

7-27

TM 55-1510-221-10

Figure 7-23. Clim

b Tw

o-Engine - Flaps 40%

7-28

TM

55-1510-221-10

Figure 7-24. Clim

b One-E

ngine Inoperative

7-2

9

TM 55-1510-221-10

Figure 7-25. Service Ceiling One-engine-Inoperative

7-30

TM 55-1510-221-10

Figure 7-26. Tim

e, Fuel, and Distance to C

ruise Clim

b

7-31

TM 55-1510-221-10

Figure 7-27. Maximum Cruise Power 1900 RPM ISA-30°C (Sheet 1 of 2)

7-32

TM 55-1510-221-10


7-33

TM 55-1510-221-10


7-34

TM 55-1510-221-10


7-35

TM 55-1510-221-10

Figure 7-29. Maximum Cruise Power 1900 RPM ISA- 10°C (Sheet 1 of 2)

7-36

TM 55-1510-221-10

Figure 7-29. Maximum Cruise Power 1900 RPM ISA- 10°C (Sheet 2 of 2)

7-37

TM 55-1510-221-10

Figure 7-30. Maximum Cruise Power 1900 RPM ISA (Sheet 1 of 2)

7-38

TM 55-1510-221-10

Figure 7-30. Maximum Cruise Power 1900 RPM ISA (Sheet 2 of 2)

7-39

TM 55-1510-221-10

Figure 7-37. Maximum Cruise Power 1900 RPM ISA+ 10°C (Sheer 1 of 2)

7-40

TM 55-1510-221-10

Figure 7-31. Maximum Cruise Power 1900 RPM ISA+ 10°C (Sheet 2 of 2)

7-41

TM 55-1510-221-10

Figure 7-32. Maximum Cruise Power 1900 RPM ISA+20°C (Sheet 1 of 2)

7-42

TM 55-1510-221-10


7-43

TM 55-1510-221-10


7-44

TM 55-1510-221-10


7-45

TM 55-1510-221-10


7-46

TM 55-1510-221-10


7-47

TM 55-1510-221-10

Figure 7-35. Maximum Cruise Speeds 1900 RPM

7-48

TM 55-1510-221-10

Figure 7-36. Maximum Cruise Power 1900 RPM

7-49

TM 55-1510-221-10

Figure 7-37. Fuel Flow At Maximum Cruise Power 1900 RPM

7-50

TM 55-1510-221-10

Figure 7-38. Range Profile -

Maxim

um C

ruise Power 1900 R

PM

7-51

TM 55-1510-221-10

Figure 7-39. Maximum Range Power 1700 RPM ISA-30°C (Sheet 1 of 2)

7-52

TM 55-1510-221-10


7-53

TM 55-1510-221-10


7-54

TM 55-1510-221-10


7-55

TM 55-1510-221-10

Figure 7-41. Maximum Range Power 1700 RPM ISA- 10°C (Sheet 1 of 2)

7-56

TM 55-1510-221-10

MAXIMUM RANGE POWER1700 RPMISA-10°C

Figure 7-41. Maximum Range Power 1700 RPM ISA- 10°C (Sheet 2 of 2)

7-57

TM 55-1510-221-10

Figure 7-42. Maximum Range Power 1700 RPM ISA (Sheet 1 of 2)

7-58

TM 55-1510-221-10

MAXIMUM RANGE POWER1700 RPM

ISA

Figure 7-42. Maximum Range Power 1700 RPM ISA (Sheet 2 of 2)

7-59

TM 55-1510-221-10


ISA +10 °C

Figure 7-43. Maximum Range Power 1700 RPM ISA+10°C (Sheet 1 of 2)

7-60

TM 55-1510-221-10


ISA+10 °C

Figure 7-43. Maximum Range Power 1700 RPM ISA+ 10°C (Sheet 2 of 2)

7-61

TM 55-1510-221-10


ISA +20 °C


7-62

TM 55-1510-221-10


ISA +20 °C


7-63

TM 55-1510-221-10


ISA+30 °C

Figure 7-45. Maximum Range Power 1700 RPM ISA+3O°C (Sheet 1 of 2)

7-64

TM 55-1510-221-10


ISA +30 °C

Figure 7-45. Maximum Range Power 1700 RPM ISA+3O°C (Sheet 2 of 2)

7-65

TM 55-1510-221-10


ISA+37 °C


7-66

TM 55-1510-221-10


ISA +37 °C


7-67

TM 55-1510-221-10

Figure 7-47. Maximum Endurance Power 1700 RPM ISA-30°C (Sheet 1 of 2)

7-68

TM 55-1510-221-10


7-69

TM 55-1510-221-10


7-70

TM 55-1510-221-10


7-71

TM 55-1510-221-10

Figure 7-49. Maximum Endurance Power 1700 RPM ISA-10° (Sheet 1 of 2)

7-72

TM 55-1510-221-10


7-73

TM 55-1510-221-10

Figure 7-50. Maximum Endurance Power 1700 RPM ISA (Sheet 1 of 2)

7-74

TM 55-1510-221-10

Figure 7-50. Maximum Endurance Power 1700 RPM ISA (Sheet 2 of 2)

7-75

TM 55-1510-221-10

Figure 7-51. Maximum Endurance Power 1700 RPM ISA+ 10°C (Sheet 1 of 2)

7-76

TM 55-1510-221-10

Figure 7-51. Maximum Endurance Power 1700 RPM ISA+ 10°C (Sheet 2 of 2)

7-77

TM 55-1510-221-10

Figure 7-52, Maximum Endurance Power 1700 RPM ISA+20°C (Sheet 1 of 2)

7-78

TM 55-1510-221-10

Figure 7-52. Maximum Endurance Power 1700 RPM ISA+20°C (Sheet 2 of 2)

7-79

TM 55-1510-221-10

Figure 7-53. Maximum Endurance Power 1700 RPM ISA+3O°C (Sheet I of 2)

7-80

TM 55-1510-221-10


7-81

TM 55-1510-221-10


7 - 8 2

TM 55-1510-221-10


7-83

TM

55-1510-221-10

7-84

Figure 7-55. R

ange Profile - L

ong Range P

ower 1700 R

PM

TM 55-1510-221-10

R A N G E P R O F I L E - 5 4 2 G A L L O N S U S A B L E F U E L

ASSOCIATED CONDITIONS:

STANDARD DAY (ISA)

EXAMPLE:

WEIGHT ............... 15,090 LBS BEFORE ENGINE START PRESSURE ALTITUDE ........................ 20.000 FTFUEL ................... AVIATION KEROSENEFUEL DENSITY RANGE AT:....... 6.7 LBS/GALWIND .............. 1059 NM.................. ZERO MAXIMUM CRUISE POWER

MAXIMUM ENDURANCE POWER ........ 1079 NMLONG RANGE POWER .................... 1107 NM

NOTE: RANGE ALLOWS FOR TAXI AND RUNUP; INCLUDES CRUISE CLIMB AND DESCENT;AND ALLOWS FOR 45 MINUTES RESERVE FUEL AT LONG RANGE POWER.

Figure 7-56. Range Profile - 542 Gallons Useable Fuel

7-85

TM

55-1510-221-10

Figure 7-57. E

ndurance Profile - 542 G

allons Useable F

uel

7-86

TM 55-1510-221-10

T I M E , F U E L , A N D D I S T A N C E T O D E S C E N D

ASSOCIATED CONDITIONS: EXAMPLE:

POWER . . . AS REQUIRED TO DESCEND INITIAL ALTITUDE. . . 26,000 FTAT 1500 FT/MIN FINAL ALTITUDE . . . . 4732 FT

LANDING GEAR.. UPFLAPS . . 0%

TIME TO DESCEND . . (17.7-3.1) 15 MINFUEL TO DESCEND . . (198.5-50.0) 149 LBSDISTANCE TO DESCEND . . (86.5-15.3) 71 NM

DESCENT SPEED: VMO/MMO

Figure 7-58. Time, Fuel, And Distance to Descend

7-87

TM

55-1510-221-10

Figure 7-59. C

limb - B

alked Landing

7-88

TM

55-1510-221-10

Figure 7-60. Norm

al Landing W

ithout propeller Reversing - Flaps 100%

7-89

TM 55-1510-221-10

L A N D I N G D I S T A N C E W I T H O U T P R O P E L L E R R E V E R S I N G - F L A P S 0 %ASSOCIATED CONDITIONS: EXAMPLE:

POWER RETARDED TO MAINTAIN APPROPRIATE FLAPS - 100% LANDINGDESCENT RATE ON FINAL APPROACH. ABOVE DISTANCE OVER13.500 LBS USE 500 FT/MIN: AT OR BELOW 50-FT OBSTACLE.. 2510 FT13,500 LBS USE 600 FT/MIN:

FLAPS 0%RUNWAY PAVED, LEVEL, DRY SURFACEAPPROACH SPEED.. IAS AS TABULATED 50-FT OBSTACLE. _. 3410 FT

APPROACH SPEED 111 KTSBRAKING MAXIMUM

LANDING WEIGHT 10,309 LBSFLAPS-UP LANDING

DISTANCE OVER

NOTES: 1. LANDING WITH FLAPS FULL DOWN (100%) IS NORMAL PROCEDURE. USE THE GRAPH BELOW WHEN ITIS NECESSARY TO LAND WITH FLAPS UP (0%).

2. TO DETERMINE FLAPS-UP LANDING DISTANCE, READ FROM THE “NORMAL LANDING DISTANCE WITH-OUT PROPELLER REVERSING - FLAPS 100%” GRAPH THE LANDING DISTANCE APPROPRIATE TOFAT, ALTITUDE, WIND, AND 50-FT OBSTACLE, THEN ENTER THE GRAPH BELOW WITH DERIVED VALUEAND READ THE FLAPS-UP LANDING DISTANCE.

Figure 7-61. Landing Distance Without Propeller Reversing - Flaps 0%

7-90

TM 55-1510-221-10

C H A P T E R 8

N o r m a l P r o c e d u r e s

Section I. MISSION PLANNING

8-l. MISSION PLANNING.

Mission planning begins when the mission isassigned and extends to the preflight check of theaircraft. It includes, but is not limited to, checks ofoperating limits and restrictions; weight, balance,and loading; performance: publications; flight plan;and crew and passenger briefings. The pilot in com-mand shall insure compliance with the contents ofthis manual that are applicable to the mission.

8-2. OPERATING LIMITS AND RESTRICTIONS.

Minimum, maximum, normal, and cautionaryoperational ranges represent careful aerodynamicand structural calculations, substantiated by flighttest data. These limitations must be adhered to dur-ing all phases of the mission. Refer to Chapter 5,OPERATING LIMITS AND RESTRICTIONS, fordetailed information.

8-3. WEIGHT, BALANCE, AND LOADING.

The aircraft must be loaded and weight and bal-ance verified per Chapter 6. WEIGHT, BALANCE,AND LOADING.

8-4. PERFORMANCE.

Refer to Chapter 7, PERFORMANCE DATA,to determine the capability of the aircraft for theentire mission. Consideration must be given tochanges in performance resulting from variation inloads, temperatures, and pressure altitudes. Recordthe data on the Performance Planning Card for usein completing the flight plan and for referencethroughout the mission.

8-5. FLIGHT PLAN.

A flight plan must beAR 95-1, DOD FLIP, and

8-6. CREW BRIEFINGS.

completed and filed perlocal regulations.

A crew briefing must be conducted for a thor-ough understanding of individual and team respon-sibilities. The briefing should include, but not belimited to, copilot, crew chief, and ground crewresponsibilities and the coordination necessary tocomplete the mission most efficiently. A review ofvisual signals is desirable when ground guides do nothave a direct voice communications link with thecrew. Refer to Section VI for crew briefings.

Section II. OPERATlNG PROCEDURES AND MANEUVERS

8-7. OPERATING PROCEDURES AND MANEU-VERS.

This section deals with normal procedures andincludes all steps necessary for safe and efficientoperation of the aircraft from the time a preflightbegins until the flight is completed and the aircraftis parked and secured. Unique feel, characteristics,and reaction of the aircraft during various phases ofoperation and the techniques and procedures usedfor taxiing, takeoff. climb. etc., are described,including precautions to be observed. Only theduties of the minimum crew necessary for the actualoperation of the aircraft are included. For operationof avionics equipment. refer to the operating hand-books that accompany the aircraft loose tools.

8-8. ADDITIONAL DATA.

Additional crew duties are covered as necessaryin Section VI, CREW DUTIES. Mission equipmentchecks are contained in Chapter 4, MISSIONEQUIPMENT. Procedures specifically related toinstrument flight that are-different from normal pro-cedures are covered in this section following normalprocedures. Descriptions of functions, operations,and effects of controls are covered in Section III,FLIGHT CHARACTERISTICS, and are repeated inthis section only when required for emphasis.Checks that must be made under adverse environ-mental conditions, such as desert and cold weatheroperations, supplement normal procedures checks inthis sect ion and are covered in Sect ion V,ADVERSE ENVIRONMENTAL CONDITIONS.

8-1

TM 55-1519-221-10

8-9. CHECKLIST.

Normal procedures are given primarily in check-list form and are amplified as necessary in accompa-nying paragraph form when a detailed description ofa procedure or maneuver is required. A condensedversion of the amplified checklist, omitting allexplanatory text, is contained in the Operator’s andCrewmember’s Checklist, TM 55-1510-221-CL. Toprovide for easier cross referencing, the proceduralsteps are numbered to coincide with the correspond-ing numbered steps in TM 55-1510-221-CL.

8-10. USE OF CHECKLIST.

Although a good working knowledge of all air-craft procedures is desirable, it is not mandatorythat they be committed to memory. The pilot isresponsible for the initiation and accomplishment ofall required checks. Checklist items will be calledout orally and the action verified using the pilot’schecklist (-CL). The copilot will normally read thechecklist and perform such duties as indicated, aswell as those directed by the pilot. ‘As required”will not be used as a response; instead the actualposition or setting of the unit or item, such as “ON”or “UP“ or “APPROACH’ will be stated. Uponcompletion of each checklist, the copilot will advisethe pilot that the checklist called for has been com-pleted.

8-11. CHECKS.

Items which apply only to night or only toinstrument flying shall have an "N" or “I” respec-tively, immediately preceding the check to which itis pertinent. The symbol “O’ shall be used to indi-cate “if installed.’ Those duties which are theresponsibility of the copilot, at the command of thepilot, will be indicated by a circle “O” around thestep number, i.e., ➃ Circuit breakers - In. The starsymbol " ★ " indicates an operational check containedin the performance section of the condensed checkl-ist. The asterisk symbol "*" indicates that perfor-mance of the step is mandatory for all thru-flights.The asterisk applies only to checks performed priorto takeoff. Placarded items appear in upper case.

8-12. BEFORE EXTERIOR CHECK.

* 1. Publications - Check DA Forms 2408-12,-13, -14, and -18, DD Form 365-4, locallyrequired forms and publications, and avail-ability of operator’s manual (-10) and check-list (-CL).

* 2. Oxygen system - Check that oxygen quantityis sufficient for the entire mission.

8-2 Change 4

If high or gusty winds are present, and theflight controls are unlocked, control sur-faces may be damaged by buffeting.

* 3. Flight controls - Unlock and check.

* 4. Parking brake - Set.

The elevator trim system must not beforced past the limits which are indicatedon the elevator trim tab position indica-tor.

5. Elevator trim - Set to “0” (neutral).

Do not cycle landing gear handle on theground.

* 6. Gear - DN.

* 7. ICC vane pull handles - In.

* 8. Keylock switch - ON.

* 9. Battery switch - ON.

10. Ice vane switches - RETRACT.

11. Lighting systems - Check as required, toinclude navigation lights, recognition lights,landing/taxi light, wing ice lights, beacons,emergency lights, and interior lights, thenOFF.

NOTE

The emergency lights override switchshould be placed in the TEST positionand the emergency lights (5) checked forillumination and intensity. A dim lightindicates a weak battery pack. At thecompletion of the check, the switch mustbe cycled from the TEST position to theOFF/RESET position and then placed inAUTO.

12. Fuel gages - Check fuel quantity and

gage operation.

TM 55-1510-221-10

★ 13. Pitot tubes (2), stall warning vane, heatedfuel fuel vents (2) - Check.

1. Stall warning heat switch - ON.

2. Pitot heat switches (2) - ON. Check coverremoved.

3. Fuel vent heat switches (2) - ON.

4. Left wing heated fuel vent - Check by feel forheat and condition.

5. Stall warning vane - Check by feel for heatand condition.

6. Right wing heated fuel vent - Check by feelfor heat and condition.

7. Stall warning heat switch - OFF.

8. Pitot heat switches (2) - OFF.

9. Heated fuel vent switches (2) - OFF.

14. Battery switch - As required.

Change 4 8-2.1/(8-2.2 blank)

TM 55-1510-221-10

1. Area 1 - Left wing, landing gear, engine, nacelle and propeller2. Area 2 - Nose section3. Area 3 - Right wing, landing gear, engine, nacelle and propeller4. Area 4 - Fuselage, right side5. Area 5 - Empennage6. Area 6 - Fuselage, left side

Figure 8-1. Exterior Inspection

AP 012082

8-3

TM 55-1510-221-10

8-13.

8-14.

15. Mission equipment and circuit breakers -Check and set.

16. Toilet - Check condition.

17. Emergency equipment - Check that allrequired emergency equipment is availableand that fire extinguishers and first aid kitshave current inspection dates.

Parachutes - Check secure and for currentinspection and repack dates.

EXTERIOR CHECK.

FUEL SAMPLE.

NOTE

Fuel and oi l quant i ty check may beperformed prior to EXTERIOR CHECK.During warm weather open fuel cap slowly toprevent being sprayed by fuel under pressuredue to thermal expansion.

* 1. Fuel sample - Check collective fuel sample froma l l d r a i n s for possible contamination.Thru-flight check is only required if aircraft hasbeen refueled.

8-15. LEFT WING, AREA I.

1. Left wing area - Check as follows (fig. 8-l):

* a. General condition - Check for skin damagesuch as buckling, splitting, distortion, dents, or fuelleaks.

b. Flaps - C h e c k f o r f u l l r e t r a c t i o n(approximately 0.25 inch play) and skin damage such asbuckling, splitting, distortion, or dents.

c. Fuel sump drains (3) - Check for leaks.

d. Controls and moveable trim tab - Checksecurity and moveable trim tab position.

NOTE

All static wicks (27) must be installed foroptimum radio performance.

e. Static wicks (4) - Check security andcondition.

f. Wing pod, navigation lights, deice boots andantennas - Check condition.

g. Recognition light - Check condition.

h. Outboard antenna set - Check condition.

i. Main tank fuel and cap - Check fuel levelvisually, condition of seal, and cap tight and properlyinstalled.

j. Outboard wing fuel vent - Check free ofobstructions.

k. Outboard deice boot - Check for securebonding, cracks, loose patches, stall strips, and generalcondition.

1. Stall warning vane - Check free.

m. Monopole antenna - Check generalcondition.

* n. Tiedown - Release.

o. Inboard dipole antenna set - Check forsecurity and cracks at mounting points. Check bondingsecure, boots free of cuts and cracks.

p. Wing ice light - Check condition.

q. AC GPU access door - Secure.

r. Recessed and heated fuel vents - Check free ofobstructions.

s. Inverter inlet and exhaust louvers - Checkfree of obstructions.

8-16. LEFT MAIN LANDING GEAR.

1. Left main landing gear - Check as follows:

* a. Tires - Check for cuts, bruises, wear, properinflation and wheel condition.

b. Brake assembly - Check brake lines fordamage or signs of leakage, brake linings for wear (0.25inch maximum, between housing and lining carrier),brake deice assembly and bleed air hose for conditionand security.

* c. Shock strut - Check for signs of leakage,minimum strut extension (5.56 inches), and that left andright strut extension is approximately equal.

d. Torque knee - Check condition.

e. Safety switch - Check condition, wire, andsecurity.

★ f. Fire extinguisher pressure - Check pressurewithin limits.

g. Wheel well, doors, and linkage - Check forsigns of leaks, broken wires, security, and generalcondition.

8-4 Change 4

h. Fuel sump drains (forward) - Check forleaks.

8-17. LEFT ENGINE AND PROPELLER.

1. Left engine - Check as follows:

A cold oil check is unreliable. Oil shouldbe checked within 10 minutes after stop-ping engine. If more than 10 minuteshave elapsed, motor engine for 30 sec-onds, then recheck. If more than 10 hourshave elapsed, run engine for 2 minutes,then recheck. Add oil as required. Do notoverfill.

* a. Engine oil - Check oil level no morethan 2 quarts low, cap secure, lockingtab aft, and access door locked.

b. Engine compartment, left side - Checkfor fuel and oil leaks, security of oilcap, door locking pin, and general con-dition.

NOTE

Secure front cowling latches first.

c.

d.

* e.

* f.

g.

h.

i.

j.

Left cowl locks - Locked.

Left exhaust stack - Check for cracksand free of obstructions.

Propeller blades and spinner - Checkblade condition, boots, security of spin-ner and free propeller rotation.

Engine air inlets and ice vane - Checkfree of obstruction and ice vaneretracted.

Bypass door - Check condition.

Right cowl locks - Locked.

Right exhaust stack - Check for cracksand free of obstructions.

Engine compartment, right side -Check for fuel and oil leaks, ice vanelinkage, door locking pin, and generalcondition. Lock compartment accessdoor.

TM 55-1510-221-10

8-18. CENTER SECTION LEFT SIDE.

1. Center section - Check as follows:

a. Heat exchanger inlet and outlet - Checkfor cracks and free of obstruction.

b. Auxiliary tank fuel sump drain - Checkfor leaks.

C. Inboard deice boot - Check for securebonding, cracks, loose patches, andgeneral condition.

* d. Auxiliary tank fuel gage and cap -Check fuel level visually, condition ofseal, and cap tight and properlyinstalled.

e. Monopole antenna - Check condition.

8-19. FUSELAGE UNDERSIDE.

1. Fuselage underside - Check as follows:

* a. General condition - Check for skindamage such as buckling, splitting, dis-tortion, dents, or fuel leaks.

b. Antennas - Check security, and generalcondition.

8-20. NOSE SECTION, AREA 2.

1. Nose section - Check as follows:

a.

b.

c.

d.

e.

f.

* g.

* h.

i.

j.

k.

Outside air temperature probe - Checkcondition.

Avionics door, left side - Check secure.

Air conditioner exhaust - Check free ofobstructions.

Wheel well - Check for signs of leaks,broken wires and general condition.

Doors and linkage - Check condition,security, and alignment.

Nose gear turning stop - Check condi-tion.

Tire - Check for cuts, bruises, wear,appearance of proper inflation, andwheel condition.

Shock strut - Check for signs of leakageand 3.0 inches minimum extension.

Torque knee - Check condition.

Shimmy damper and linkage - Checkfor security and condition.

Landing and taxi lights - Check forsecurity and condition.

8-5

TM 55-1510-221-10

l.

m.

n.

o.

p.

q .

r.

Pitot tubesalignment,

- Check covers removed,security, and free of

obstructions.

Radome - Check condition.

Antenna podtion.

- Check general condi-

Monopole antennas (2) - Check generalcondition.

Windshields and wipers - Check wind-shield for cracks and cleanliness andwipers for contact with glass surface.

Air conditioner inlet - Check free ofobstructions.

Avionics door, right side - Checksecure.

8-21. CENTER SECTION, RIGHT SIDE.

1. Center section - Check as follows:

a.

b.

c.

* d.

e.

f.

g.

h.

i.

j.

Inboard deice boot - Check for securebonding, cracks, loose patches, andgeneral condition.

Battery access panel - Secure.

Battery vents - Check free of obstruc-tion.

Auxiliary tank fuel gage and cap -Check fuel level visually, condition ofseal, and cap tight and properlyinstalled (locking tab aft).

Battery compartment drain - Checkfree of obstruction.

Battery ram air intake - Check free ofobstruction.

INS TAS temperature probe - Checkcondition and free of obstructions.

Auxiliary tank fuel sump drain - Checkfor leaks.

Heat exchanger inlet and outlet - Checkfor cracks and free of obstructions.

Monopole antenna - Check condition.

8-22. RIGHT ENGINE AND PROPELLER.

1. Right engine and propeller - Check as fol-lows:

A cold oil check is unreliable. Oil shouldbe checked within 10 minutes after stop-

ping engine. If more than 10 minuteshave elapsed, motor engine for 30 sec-onds, then recheck. If more than 10 hourshave elapsed, run engine for 2 minutes,then recheck. Add oil as required. Do notoverfill.

* a.

b.

c.

d.

* e.

* f.

g.

h.

i.

j.

Engine oil - Check oil level, oil capsecure (locking tab aft), and accessdoor locked.

Engine compartment, left side - Checkfor fuel and oil leaks, security of oilcap, door locking pins, and generalcondition.

Left cowl locks - Locked.

Left exhaust stack - Check for cracksand free of obstructions.

Propeller blades and spinner - Checkblade condition, boots, security of spin-ner, and free propeller rotation.

Engine air inlets and ice vane - Checkfree of obstruction and ice vaneretracted.

Bypass door - Check condition.

Right cowl locks - Locked.

Right exhaust stack - Check for cracksand free of obstructions.

Engine compartment, right side -Check for fuel and oil leaks, ice vanelinkage, door locking pins, and generalcondition. Lock compartment accessdoor.

8-23. RIGHT MAIN LANDING GEAR.

1. Right main landing gear - Check as follows:

a.

* b.

c.

* d.

Fuel sump drains (forward) - Check forleaks.

Tires - Check for cuts, bruises, wear,proper inflation and wheel condition.

Brake assembly - Check brake lines fordamage or signs of leakage, brake lin-ings for wear (0.25 inch maximum,between piston housing and lining car-rier), brake deice assembly and bleedair hose for condition and security.

Shock strut - Check for signs of leak-age, minimum strut extension (5.50inches), and that left and right strutextension is approximately equal.

8-6

e.

f.

★ g.

h.

Torque knee - Check condition.

Safety switch - Check condition, wire, andsecurity.

Fire extinguisher pressure - Checkpressure within limits.

Wheel well, doors, and linkage - Checkfor signs of leaks, broken wires, security,and general condition.

8-24. RIGHT WING, AREA 3.

1. Right wing - Check as follows:

a.

b.

c.

d.

e.

f.

* g.

* h.

i.

j.

k.

1.

m.

n.

o.

p.

Recessed and heated fuel vents - Checkfree of obstructions.

Inverter inlet and exhaust louvers - Checkfree of obstructions.

DC GPU access door - Secure.

Inboard dipole antenna set - Check forsecurity and cracks at mounting points,bonding secure, free of cuts and cracks.

Wing ice light - Check condition.

Outboard deice boot - Check for securebonding, cracks, loose patches, stall strips,and general condition.

Tiedown - Release.

Main tank fuel and cap - Check fuel levelvisually, condition of seal, and cap tightand properly installed.

Outboard wing fuel vent - Check free ofobstructions.

Outboard antenna set - Check condition.

Recognition light - Check condition.

Wing pod navigation lights, deice bootsand antennas - Check condition.

Static wicks (4) - Check security andcondition.

Controls - Check security and condition ofground adjustable tab.

Fuel sump drains (3) - Check for leaks.

Flaps - Check for full retraction(approximately 0.25 inch play) and skindamage, such as buckling, splitting,distortion, or dents.

q . Chaff dispenser - Check number of chaffsin payload module and for security.

* r.

TM 55-1510-221-10

General condition - Check for skindamage such as buckling, splitting,distortion, dents, or fuel leaks.

8-25. FUSELAGE RIGHT SIDE, AREA 4.

1. Fuselage right side - Check as follows:

* a.

b.

C.

d.

e.

f.

g.

h.

i.

j.

k.

l.

General condition - Check for skindamage such as buckling, splitting,distortion or dents.

Emergency light - Check condition.

Flare/chaff dispenser - Check number offlares in payload module and for security.

Beacon - Check condition.

Underside fuselage antennas - Cheekgeneral condition.

Towel bar antennas (2) - Check general

P-band antenna - Check general condition.

Tailcone access door - Check secure.

Oxygen filler door - Check secure.

Static ports - Check clear of obstructions.

APR 44 antennas (2) - Check.

Emergency locator transmitter antenna -Check condition.

8-26. EMPENNAGE, AREA 5.

1. Empennage - Check as follows:

a.

b.

c.

d.

e.

Vertical stabilizer, rudder, and trim tab -Check for skin damage, such as buckling,distortion, or dents, and trim tab rig.

Static wicks (19) - Check installed.

Antennas - Check security, and generalcondition.

Deice boots - Check for secure bonding,cracks, loose patches, and generalcondition.

Horizontal stabilizer, and elevator - Checkfor skin damage, such as buckling,distortion and dents.

NOTE

Any difference between the indicated positionon the trim tab position indicator and the actual

Change 4 8-7

TM 55-1510-221-10

position of the elevator trim tab signifies anunairworthy condition and must be correctedprior to the next flight of the aircraft.

f. Elevator trim tab -position.

WARNING

Verify “0” (neutral)

If the possibility of ice accumulation on thehorizontal stabilizer or elevator exists, takeoffwill not be attempted.

g .

h.

i .

Position and beacon lights - Checkcondition.

Rotating boom dipole antenna - Checkcondition and positron.

Wide band data link antenna pod - Checkfor cracks and chips.

8-27. FUSELAGE, LEFT SIDE, AREA 6.

1. Fuselage left side - Check as follows:

* a.

b.

c.

d.

e.

f.

g.

h.

i.

* j.

General condition - Check for skindamage such as buckling, distortion, ordents.

Static ports - Check clear of obstructions.

ELT-ARMED.

APR-44 antennas (2) - Check.

P band antenna - Check general condition.

Towel bar antennas (2) - Check generalcondition.

Emergency light - Check condition.

Cabin door - Check door seal and generalcondition.

Fuselage to side - Check generalcondition and antennas.

Chocks and tiedowns - Check removed.

8-28. INTERIOR CHECK.

1. Cargo/loose equipment - Check secure.

2. Cabin/cargo doors - Test and lock:

a. Cabin door - Check closed and latched bythe following.

(1) Safe arm and diaphram plunger -Check position (lift door step).

(2) Index marks on rotary cam locks (6)- Check aligned with indicatorwindows.

b. Cargo door - Check closed and latched bythe following:

(1)

(2)

(3)

(4)

(5)

(6)

(7)

(8)

(9)

Upper handle - Check closed andlatched. (Observe through cargo doorlatch handle access cover window.)

Index marks on rotary cam locks (4)- Check aligned with indicatorwindows.

Lower pin latch handle - Checkclosed and latched. (Observe throughcargo door lower latch handle access ,cover window.)

Carrier rod - Check indicator alignedwith orange stripe on carrier rod.(Observe through window aft lowercomer.)

Battery switch - OFF.

Cargo door - Check closed andlatched.

Cabin door - Close but leaveunlatched. Check CABIN DOORannunciator light illuminated.

Cabin door - Open. Check CABINannunciator light

extinguished

Battery switch - ON. Check CABINDOOR annunciator light illuminated.

(10) Cabin door - Close and latch. CheckCABIN DOOR annunciator lightextinguished.

(11) Battery switch - OFF.

3. Emergency exit - Check secure and keyremoved.

4. Mission cooling ducts - Check open and free ofobstructions.

5. Flare/chaff dispenser preflight test -Completed.

6. KKY-28/58 key loaded - As required.

★ 7. Crew briefing - As required. Refer to SectionVI.

8-8 Change 4

TM 55-1510-221-10

8-29. BEFORE STARTING ENGINES.

★ 1. Oxygen system - Check as required.

a .

b .

c .

d .

e.

f.

g.

Oxygen supply pressure gages - Check.

Supply control lever (green) - ON.

Diluter control lever - 100% OXYGEN.

Emergency control lever (red) - Set toTEST MASK position while holding maskdirectly away from face, then return toNORMAL.

Oxygen masks - Put on and adjust.

Emergency pressure control lever - Set toTEST MASK position and check mask forleaks, then return lever to NORMAL.

Flow indicator - Check, during inhalationblinker appears, during exhalation blinkerdisappears). Repeat a minimum of 3times.

2. Circuit breakers - Check in.

* 3. Overhead control panel switches - Set as

a.

b.

c.

d.

e.

f.

g.

h.

i.

j.

k.

1.

Light dimming controls - As required.

Cabin temperature mode selector switch -OFF.

Ice & rain switches - As required.

Exterior light switches - As required.

Master panel lights switch - As required.

Inverter switches - As required.

Avionics master power switch - Asrequired.

Environmental switches - As required.

Autofeather switches - OFF.

# 1 engine start switches - OFF.

Master switch - As required.

#2 engine start switches - OFF.

* 4. Fuel panel switches - Check as follows:

a. Standby fuel pump switches - OFF.

b. Auxiliary transfer override switches -AUTO.

c. Crossfeed switch - OFF.

5. Magnetic compass - Check for fluid, headingand current deviation card.

* 6. Pedestal controls - Set as follows:

CAUTION

Movement of power levers into reverse rangewhile engines are shut down may result inbending and damage to control linkages.

a. POWER levers - IDLE.

b. Propeller levers - HIGH RPM.

c. CONDITION levers - FUEL CUTOFF.

d. Flaps- UP.

e. Friction locks - Check and set.

* 7. Pedestal extension switches - Set as follows:

a. Flare/chaff dispenser control - SAFE.

b. Avionics - As required.

c. Rudder boost switch - ON.

8. Gear alternate engage and ratchet handles -stowed.

9. Outside air temperature gage - Check, notecurrent reading.

10. Instrument panel - Check and set as follows:

a.

b.

c.

d.

e.

f.

g.

h.

i.

Pilot’s and copilot’s course indicatorswitches -As required.

Pilot’s and copilot’s RMI switches - Asrequired.

Pilot’s and copilot’s microphone switch -As required.

Pilot’s and copilot’s compass switches -As required.

Gyro switches - SLAVE.

Flight instruments - Check instruments forprotective glass, warning flags (12 pilot, 6 copilot), static readings, and headingcorrection card.

Radar - OFF.

APR-39 and APR-44 - OFF.

Engine instruments - Check for protectiveglass and static readings.

Change 4 8-9

TM 55-1510-221-10

11. Deleted.

12. Mission panel switches and circuit breakers -set and OFF.

13. Pressurization controls - Set.

14. Subpanels - Check and set as follows:

a

b.

c.

d.

Fire protection test switch - OFF.

Landing, taxi, and recognition lights -OFF.

Landing gear control switch - RecheckDN.

Cabin lights - As Required.

15. Pilot’s static air source - NORMAL.

NOTE

Do not use alternate static source duringtakeoff and landing except in an emergency.Pilot’s instruments will show a variation inairspeed and altitude.

16. Pilot’s and copilot’s audio control panels - Asrequired.

17. Deleted.

★ 18. Fuel pumps/crossfeed operation - Check as

a .

b .

c .

d .

e.

f.

a .

b .g.

h.

i.

Fire pull handles - Pull.

Standby fuel pump switches - On.

Battery switch - ON.

#1 fuel pressure and #2 fuel pressurewarning lights - Illuminated

Fire pull handles - In.

#l fuel press and #2 fuel press warningannunciator lights - Extinguished.

Standby fuel pump switches - Off.

#1 fuel pressure and #2 fuel pressurewarning lights - Illuminated.

Crossfeed - check. Check systemoperation by activating switchmomentarily left then right, noting that #1FUEL PRESS and #2 FUEL PRESSwarning annunciator lights extinguish andthat the FUEL CROSSPEED advisoryannunciator light illuminates as switch isenergized.

19. AC and DC GPU - As required.

8-10 Change 4

20. External power advisory annunicator lights -As required. (Aircraft EXTERNAL POWERand mission EXT DC PWR ON annunciatorlights illuminated.)

21.

★ 22.

DC power - Check. (22 VDC minimum forbattery, 28 maximum for GPU starts).

Annunciator panels - Test as required.

a MASTER CAUTION, MASTERWARNING, #1 FUEL PRESS, #2 FUELPRESS, GEAR DN, L BL AIR FAIL, RBL AIR FAIL, INST AC, #1 DC GEN. #1INVERTER, #1 NO FUEL XFR, #2 NOFUEL XFR, #2 INVERTER, #2 DC GEN,- Check illuminated.

b. ANNUNCIATOR TEST switch - Pressand hold. Check that the annunciator

panels, FIRE PULL handle lighbeacon lights, ANT Azimuth

ts, marker,indicator,

MASTER CAUTION and MASTERWARNING lights are illuminated. Releaseswitch and check that all lights exceptthose in step (1) are extinguished.

c. MASTER CAUTION and MASTERWARNING lights - Press. Check that bothlights extinguish.

★ 23. Stall and gear warning system - Check asfollows:

a. STALL WARN TEST switch - TEST.Check that warning horn sounds.

b. LDG GEAR WARN TEST switch -TEST. Check that waming horn soundsand that the LDG GEAR CONTR handlewarning lights (2) illuminate.

★ 24. Fire Protection system - Check as follows:

Fire Detector Test switch - Rotatecounterclockwise to check three DETRpositions. FIRE PULL handles shouldIlluminate in each position. ResetMASTER WARNING in each position.

Fire Detector Test switch - Rotatecounterclockwise to check two EXTGHpositions. SQUIB OK light, associated #1EXTGH DISCH and #2 EXTGH DISCHannunciator caution light and MASTERCAUTION LIGHT should illuminate ineach position.

25. INS - Align as required.

8-30. * FIRST ENGINE START (BATTERYSTART).

TM 55-1510-221-10

NOTE

The engines must not be started untilafter the INS is placed into the NAVmode or OFF as required.

Starting procedures are identical for bothengines. When making a battery start, the rightengine should be started first. When making aground power unit (GPU) start, the left engineshould be started first due to the GPU receptaclebeing located adjacent to the right engine. A crew-member should monitor the outside observerthroughout the engines start.

1. Avionics master switch - As required.

2. Exterior light switches - As required.

3. Propeller - Clear.

4. Ignition and engine start switch - ON. Pro-peller should begin to rotate and associated#1 IGN ON or #2 IGN ON annunciatorlight should illuminate. Associated #1 FUELPRESS or #2 FUEL PRESS light shouldextinguish.

If ignition does not occur within 10 sec-onds after moving condition lever toLOW IDLE, initiate engine clearing pro-cedure. If for any reason a startingattempt is discontinued, the entire start-ing sequence must be repeated afterallowing the engine to come to a completestop (5 minute minimum).

5. CONDITION lever (after N1 RPM stabi-lizes, 12% minimum) - LOW IDLE.

Monitor TGT to avoid a hot start. Ifthere is a rapid rise in TGT, be preparedto abort the start before limits areexceeded. During starting, the maximumallowable TGT is 1000°C for five seconds.If this limit is exceeded, use ABORTSTART procedure and discontinue start.Enter the peak temperature and durationon DA Form 2408-13.

6. TGT and N1- Monitor (TGT 1000°C maxi-mum, N1 52% minimum).

7. Oil pressure - Check (60 PSI minimum).

8. Ignition and engine start switch - OFF, after50% N1.

9. CONDITION lever - HI IDLE. MonitorTGT as the condition lever is advanced.

10. Generator switch - RESET, then ON.

8-31. SECOND ENGINE START (BATTERYSTART).

1. First engine generator load 50% or less.

2. Propeller - Clear.

3. Ignition and engine start switch - ON. Pro-peller should begin to rotate and associated#l IGN ON or #2 IGN ON annunciatorlight should illuminate. Associated #l FUELPRESS or # 2 FUEL PRESS annunciatorlight should extinguish.

If ignition does not occur within 10 sec-onds after moving condition lever toLOW IDLE, initiate engine clearing pro-cedure. If for any reason a startingattempt is discontinued, the entire start-ing sequence must be repeated afterallowing the engine to come to a completestop (5 minute minimum).

4. CONDITION lever (after N1 RPM passes12% minimum) - LOW IDLE.

Monitor TGT to avoid a hot start. Ifthere is a rapid rise in TGT, be preparedto abort the start before limits areexceeded. During starting, the maximumallowable TGT is 1000°C for five seconds.If this limit is exceeded, use ABORTSTART procedure and discontinue start.Enter the peak temperature and durationon DA Form 2408-13.

5. TGT and N1 - Monitor (TGT 1000°C maxi-mum, N1 52% minimum).

6. Oil pressure - Check (60 PSI minimum).

7. Ignition and engine start switch - OFF after50% N1.

8-11

TM 55-1510-221-10

8.

9.

10.

11.

Battery charge light - Check (light shouldilluminate approximately 6 seconds aftergenerator is brought on line. Light shouldextinguish within 5 minutes following a nor-mal engine start on battery).

Inverter switches - ON, check INVERTERannunciator lights extinguished.

Second engine generator - RESET, then ON.

CONDITION levers - As required.

8-32.

1.

2.

ABORT START.

CONDITION lever - FUEL CUTOFF.

Ignition and engine start switch - STARTERONLY.

3. TGT - Monitor for drop in temperature.

4. Ignition and engine start switch - OFF.

8-33.

1.

2.

ENGINE CLEARING.

CONDITION lever - FUEL CUTOFF.

Ignition and engine start switch - OFF (5minute minimum).

Do not exceed starter limitation of 30 sec-onds ON and 5 minutes OFF for twostarting attempts and engine clearing pro-cedure. Allow 30 minutes off before addi-tional starter operation.

3. Ignition and engine start switch - STARTERONLY (15 seconds minimum, 30 secondsmaximum).

4. Ignition and engine start switch - OFF.

8-34. * FIRST ENGINE START (GPU START).

1. INS - As required.

NOTE

The engines must not be started untilafter the INS is placed into the NAVmode or OFF as required.

2.

3.

4.

5.

6.

Avionics master switch - As required.

Exterior light switches - As required.

Propeller - Clear.

Ignition and engine start switch - ON. Pro-peller should begin to rotate and associated#1 IGN ON or #2 IGN ON should illumi-nate. Associated #1 FUEL PRESS or #2FUEL PRESS warning annunciator lightshould extinguish.

CONDITION lever (after N1, RPM stabi-lizes, 12% minimum) - LOW IDLE.

Monitor TGT to avoid a hot start. Ifthere is a rapid rise in TGT, be preparedto abort the start before limits areexceeded. During engine start, the maxi-mum allowable TGT is 1000°C for fiveseconds. If this limit is exceeded, useABORT START procedure and discon-tinue start. Enter the peak temperatureand duration on DA Form 2408- 13.

7.

8.

9.

10.

11.

12.

8-35.

1.

2.

TGT and N1 - Monitor (TGT 1000°C maxi-mum, N1 52% minimum).

Oil pressure - Check (60 PSI minimum).

Ignition and engine start switch - OFF after50% N1.

CONDITION lever - HI IDLE. MonitorTGT as the condition lever is advanced.

DC GPU - Disconnect as required,

Generator switch (GPU disconnected) -RESET, then ON.

SECOND ENGINE START (GPU START).

Propeller - Clear.

Ignition and engine start switch - ON. Pro-peller should begin to rotate and associated#1 IGN ON or #2 IGN ON annunciatorlight should illuminate. Associated #1 FUELPRESS or #2 FUEL PRESS annunciatorlight should extinguish.

8-12

TM 55-1510-221-10

CAUTION

If ignition does not occur within 10 secondsafter moving condition lever to LOW IDLE,initiate engine clearing procedure. If for anyreason a starting attempt is discontinued, theentire starting sequence must be repeated afterallowing the engine to come to a complete stop(5 minute minimum).

3. CONDITION lever (after N1 RPM passes 12%minimum) - LOW IDLE.

CAUTION NOTE

Monitor TGT to avoid a hot start. If there is arapid rise in TGT, be prepared to abort the startbefore limits are exceeded. During engine start,the maximum allowable TGT is 1000°C forfive seconds. If this limit is exceeded, useABORT START procedure and discontinuestart. Enter the peak temperature and durationon DA Form 2408-13.

For maximum cooling on the ground, turn thebleed air valve switches to ENVIRO OFFposition.

★ * 3. AC/DC power - Check for:

a. AC frequency - 394 to 406 Hz.

b. AC voltage - 104 to 124 VAC.

c. DC load - Check.4. TGT and N1 - Monitor (TGT 1000°C

maximum, N1 52% minimum).

5.

6.

Oil pressure - Check (60 PSI minimum).

Ignition and engine start switch - OFF, after50% N1

7.

8.

Propeller levers - FEATHER.

GPU - Disconnect. (Check aircraft externalpower and mission external power lightextinguished.)

9. Propeller levers - HIGH RPM.

10. Aircraft inverter switches - ON, check #1INVERTER and #2 INVERTER annunciatorlights extinguished.

11. Generator switches - RESET, then ON.

12. CONDITION levers - As required.

8-36. BEFORE TAXING.

* 1. Brake deice - As required. To activate thebrake deice system proceed as follows:

a. Bleed air valves - OPEN.

b. CONDITION levers - HI IDLE.

c. Brake deice switch - DEICE. CheckBRAKE DEICE ON advisory annunciatorlight illuminated.

* 2. Cabin temperature and mode - Set.

CAUTION

Verify airflow is present from aft cockpiteyeball outlets to insure sufficient cooling formission equipment.

d. DC voltage - 27.0 to 28.5 VDC.

WARNING

Do not operate radar in congested areas. Injurycould result to personnel in close proximity tooperating radar.

CAUTION

Do not operate the weather radar in an areawhere the nearest effective surface is 50 feet orless from the antenna reflector. Scanning suchsurfaces within 50 feet of the antenna reflectormay damage receiver crystals.

* 4. Avionics master switch - ON.

5. Mission panel - Set and checked as required.

★ 6. Automatic flight control system - Check asfollows:

a. Altitude alert.

8-13

TM 55-1510-221-10

NOTE

Pause a few seconds between each step toallow time for the proper indications.

(1)

(2)

(3)

(4)

(5)

Set alert controller more than 1000feet above altitude indicated onpilot’s altimeter. The pilot’s altimeteralert light should be extinguished.

(a)

(b)Decrease the alert controller towithin 1000 feet of the pilot’saltimeter setting. The alert lightshould illuminate.

Decrease the controller to less than250 feet above the pilot’s altimetersetting. The alert light shouldextinguish.

Increase the controller to 300 ±50feet above the pilot’s altimeterindication and check that the alertlight illuminates.

Set the desired altitude.

b. Autopilot.

(1) Autopilot controller UP TRIM, DNTRIM annunciators - CHECK notilluminated.

CAUTION

A steady illumination of UP TRIM or DNTRIM annunciator indicates that the automaticsynchronization is not functioning and theautopilot should not be engaged.

(2) Turn knob - Center.

(3) Elevator trim control switch - ON.

(4) Control wheel - Hold to mid travel.

(5) AP button - Press. AP ENGAGE andYD ENGAGE annunciators on.

(6) Deleted.

(a) Deleted.

(b) Deleted.

(7) Elevator trim follow-up - Check.

Control wheel - Hold aft of midtravel. Trim wheel should runnose down after approximately 3seconds. Trim down annunciatorshould illuminate afterapproximately 8 seconds.

Control wheel - Hold forward ofmid travel. Trim wheel shouldrun nose up after approximately 3seconds, trim up annunciatorshould illuminate afterapproximately 8 seconds, and APTRIM FAIL annunciator andMASTER WARNING lightss h o u l d i l l u m i n a t e a f t e rapproximately 15 seconds.

The elevator trim system must not be forcedbeyond the limits which are indicated on theelevator trim tab indicator.

(8) AP/YD & TRIM DISC Button -Depress through second level.Autopilot and yaw damper shoulddisengage and ELECT TRIM O F Fannunciator should illuminate. APENG and YD ENG annunciators oninstrument panel should flash 5times.

(9) Elevator trim control switch - OFF,then ON. (ELEC TRIM OFFannunciator should extinguish).

(10) AP - Re-engage.

(11) Turn controller - Check that controlwheel follows in each applieddirection, then center.

8-14 Change 4

(12)

(13)

(14)

Pitch wheel - Check that trim respondsto pitch wheel movement. (UP TRIMand DN TRIM annunciators mayilluminate).

Heading bug - Center and engage HDG.Check that control follows a turn in eachdirection.

Disengage AP by selecting GA. Checkthat AP disengages and FD commands7° nose up, wings level attitude. YDdisengage - Autopilot mode selector -STBY.

7. Electric elevator trim - Check.

a.

b.operation.

Elevator trim switch - ON.

Pilot and copilot trim switches - Check

WARNING

Operation of the electric trim system shouldoccur only by movement of pairs of switches.Any movement of the elevator trim wheelwhile depressing only one switch elementdenotes a system malfunction. The electricelevator trim control switch must then beturned OFF and flight conducted byoperating the elevator trim wheel manually.Do not use autopilot.

C.

d.

(1) Pilot and copilot. Check individualelement for no movement of trim, thencheck proper operation of both elements.

(2) Check pilot switches override copilotswitches while trimming in oppositedirections, and trim moves in directioncommanded by pilot.

Check pilot and copilot trim disconnectswhile activating trim.

Elevator trim switch - OFF then ON(ELECT T R I M O F F annunciatorextinguishes).

8. Avionics - Check and set as required.

9. INS - NAV mode, if on.

10. Flaps - Check.

11. Altimeters - Set and check.

TM 55-1510-221-10

8-37. TAXIING.

CAUTION

Excessive use of brakes with the increasedweight of this aircraft will increase thepossibility of brake failure and/or brake fire.Judicious use of the brakes is recommendedwith coordinated use of beta range.

Taxi speed can be effectively controlled by the use ofpower application and the use of the variable pitchpropellers in beta range. Normal turns may be madewith the steerable nose wheel; however, a turn may betightened by using full rudder and inside brake asnecessary. Turns should not be started with brakesalone, nor should the aircraft be pivoted sharply on onemain gear.

1. Brakes - Check.

2. Flight instruments - Check for normal operation.

8-38. ENGINE RUNUP.

1. Mission control panel - Set.

*2. Propeller manual feathering - Check as follows:

a.

b.

CONDITION lever - LOW IDLE.

Left propeller lever - FEATHER Check thatpropeller feathers.

c. Left propeller lever - HIGH RPM.

d. Repeat procedure for right propeller.

*3. Autofeather- Check as follows:

a.

b.

c.

d.

CONDITION levers - LOW IDLE.

AUTOFEATHER switch - Hold to TEST.(#1 AUTOFEATHER and #2AUTOFEATHER advisory annunciatorlights should remain extinguished.)

P O W E R l e v e r s - Advance untilAUTOFEATHER lights are illuminated(approximately 22% torque).

Left POWER lever - Retard and check:

(1)

(2)

At 16 - 21% torque - #2AUTOFEATHER light out.

At 9 - 14% torque - BothAUTOFEATHER lights out and leftpropeller starts to feather.

Change 4 8-15

TM 55-1510-221-10

e. Left POWER lever - Approximately 22%torque.

f. Repeat steps b and d for right engine.

g. POWER levers - IDLE (both lights out,neither propeller feathers).

*4. Overspeed governors - Check as follows:

a.

b.

c.

d.

e.

POWER levers - Set approximately 1950RPM (both engines).

#1 propeller governor test switch - Hold toTEST position.

#1 propeller RPM 1830 to 1910 - Check.

Repeat steps b and c for # 2 engine.

POWER levers - Set 1800 RPM.

*5. Primary governors - Check as follows:

a. POWER levers - Set 1800 RPM.

b. Propeller levers - Move aft to detent. Checkthat propeller RPM drops toRPM.

c. Propeller levers - HIGH RPM.

*6. Ice vanes - Check as follows:

a. Ice vane switches - EXTEND.

1600 to 1640

Verify torquedrop, TGT increase, and illumination of #1ICE VANE EXTEND and #2 ICE VANEEXT annunciators.

b. Ice vane switches - RETRACT. Verify returnto original torque and TGT, and that #1 ICEVANE EXTEND and #2 ICE VANE EXTannunciators extinguish.

7. CONDITION levers - HI IDLE.

8. POWER levers - IDLE.

*9. Anti-ice and deice systems - Check as follows:

a. Windshield anti-ice switches - NORMAL andHI. Check PILOT and COPILOT(individually) for loadmeter rise, then OFF.

b. Propeller switches - INNER and OUTER(momentarily). Check for loadmeter rise.

8-16 Change 4

c.

d.

e.

f.

g.

h.

Surface deice switch - SINGLE CYCLEAUTO. Check for a drop in pneumaticpressure and wing deice boot inflation andafter 6 seconds for a second drop inpneumatic pressure.

Surface deice switch - MANUAL. Check thatsurface boots inflate, and remain inflated,then OFF.

Antenna deice switch - SINGLE. Check for adrop in pneumatic pressure and antennadeice boot inflation.

Antenna deice switch - MANUAL. Check thatboots inflate, and remain inflated, then OFF.

Engine inlet lip heat switches - ON. Checkthat #1 LIP HEAT ON and #2 LIP HEAT ONadvisory annunciator lights are illuminated,and the #1 LIP HEAT and #2 LIP HEATcaution annunciator lights are extinguished,then OFF.

RADOME ANTI-ICE switch - ON. Checkthat RADOME HEAT annunciator isilluminated, then OFF.

*10. Pneumatic pressure - Check as follows:

a. Left bleed air valve switch - PNEU &ENVIRO OFF.

b. Pneumatic pressure - Check 12 to 20 PSI.

c. Right pneumatic and environmental switch -PNEU & ENVIRO OFF. Check that L BLAIR FAIL and R BLAIR FAIL, and L BL AIROFF and R BL AIR OFF annunciator lightsare illuminated.

d. Pneumatic pressure - Verify zero.

TM 55-1510-221-10

e. Left pneumatic and environmental switches -OPEN. Check that L BL AIR FAIL and R BLAIR FAIL, and L BL AIR OFF and R BL AIROFF annunciator lights are extinguished.

f. Pneumatic pressure - Verify 12 to 20 PSI.

g. Right pneumatic and environmental switches -OPEN.

* 11. Pressurization system - Check as follows:

a. Cabin door caution light - Checkextinguished.

b. Storm windows - Check closed.

c. Bleed air valve switches - Check OPEN.

d. Cabin altitude - Set 500 feet lower thanairfield elevation.

e. Cabin pressure/dump switch - TEST (hold).

f. Cabin rate-of-climb gage - Check for descend-ing indication and, when confirmed, releasecabin pressure/dump switch from TEST.

g. Aircraft altitude - Set to planned cruisealtitude plus 500 feet. (If this setting does notresult in a CABIN ALT indication of at least500 feet over takeoff field pressure altitude,adjust as required).

12. CONDITION levers - As required

13. ANTI-ICE - As required.

NOTE

If windshield anti-ice is needed prior to takeoff,use normal setting for a minimum of 15minutes prior to selecting high temperature toprovide adequate preheating and minimizeeffects of thermal shock.

8-39. BEFORE TAKEOFF.

01. Autofeather switch - ARM.

2. Bleed air valves - As required.

03. Ice & rain switches - As required.

4. Fuel panel - Check fuel quantity and switchpositions.

05. Flight and engine instruments - Check for normalindications.

06. Cabin altitude and rate-of-climb controller - Set.

07. Annunciator panels - Check (note indications).


9. Flaps - As required.

10. Trim - Set.

11. Avionics - Set.

12. Flight controls - Check

013. Departure briefing - Complete.

8-40. LINE UP.

01. Transponder - As required.

02. Engine autoignition switch - ARM.

3. Power stabilized - Check approximately 25%minimum.

04. CONDITION levers - LOW IDLE.

5. Lights - As required.

6. Mission control panel - Set.

8-41. TAKEOFF.

To aid in planning the takeoff and to obtainmaximum aircraft performance, make full use of theinformation affecting takeoff shown in Chapter 7.The data shown is achieved by setting brakes, set-ting takeoff power, and then releasing brakes. Whenrunway lengths permit, the normal takeoff may bemodified by starting the takeoff after power hasbeen stabilized at approximately 25% torque, thenapplying power smoothly so as to attain full powerno later than 65 KIAS. This will result in asmoother takeoff but will significantly increasetakeoff distance.

a. Normal Takeoff. After LINE Up procedureshave been completed, release brakes and smoothlyapply power to within 5% of target. Power shouldbe applied at a rate that will produce takeoff powerby 40 KIAS. Maintain directional control withnosewheel steering rudder, and differential power,while maintaining wings level with ailerons. Thepilot should retain a light hold on the power leversthroughout the takeoff and be ready to initiateABORT procedures if required. The copilot shouldinsure that the AUTOFEATHER advisory lights areilluminated (if applicable), adjust and maintain powerat the exact takeoff power settings, and monitor allengine instruments. The pilot will rotate at therecommended rotation speed (Vr) and establish theclimb attitude (9° to 16°) that will attain best rate-of-climb airspeed (Vy) during the initial climb.Rotation should be at a rate that will allow liftoffat liftoff airspeed (V1of).

b. Crosswind Takeoff. Position the aileron controlinto the wind at the start of the takeoff roll tomaintain a wings level attitude. Under strongcrosswind conditions, leading with upwind power atthe beginning of the takeoff roll will assist in main-taining directional control. As the nosewheel comesoff the ground, the rudder is used as necessary to

Change 3 8-17

TM 55-1510-221-10

prevent turning (crabbing) into the wind. Rotate ina positive manner to keep from side-skipping asweight is lifted from the shock struts. To preventdamage to the landing gear, in the event that the air-craft were to settle back onto the runway, remain in“slipping” flight until well clear of the ground, thencrab into the wind to continue a straight flight path.

c. Minimum Run Takeoff.

Spectacular takeoff performance can beobtained by lifting off at speeds belowthose recommended in Chapter 7. How-ever, control of the aircraft will be lost ifan engine failure occurs immediately fol-lowing liftoff until a safe speed can beattained. Except during soft field takeoff,liftoff below recommended speeds willnot be performed.

Minimum run takeoffs are performed with flapsextended to 40% although at some conditions, use offlaps during takeoff may result in the inability toattain positive single-engine climb if an engine failsimmediately after liftoff.

To compensate for torque effect during thebeginning of the takeoff roll, align the aircraft withthe nose approximately 10° right of centerline. AfterLINE UP procedures have been completed, holdbrakes firmly and apply TAKEOFF POWER, allow-ing for some increase in power as airspeed increasesduring the takeoff roll. Copilot action is the same asfor normal takeoff. Release brakes and maintaindirectional control and nosewheel steering and rud-der. Do not use brakes unless absolutely necessary.Hold the elevators in a neutral position, maintainingwings level with ailerons. Allow the aircraft to rollwith its full weight on the wheels until the recom-mended rotation speed (V,) is reached. At this speedrotate smoothly and firmly at a rate that will allowliftoff at liftoff air speed (V1of). When flight isassured, retract the landing gear.

d. Obstacle Clearance Climb. Follow proce-dures as outlined for a minimum run takeoff, to thepoint of actual liftoff. When flight is assured, retractthe landing gear and establish a wings level climbattitude, maintaining the computed obstacle clear-ance airspeed (V,). Climb at this speed until clear ofthe obstacle. After the obstacle is cleared, lower thenose slowly and accelerate to best rate-of-climb air-speed (V). Retract flaps after attaining single enginebest rate-of-climb airspeed (V yse).

8-18

NOTE

The best angle-of-climb speed (V,) is veryclose to single engine power-off stallspeed. To provide for a margin of safetyin the event of engine failure immediatelyafter takeoff, the obstacle clearance air-speed value is used in lieu of true Vx, formaximum angle takeoff climbs. Takeoffperformance data shown in Chapter 7 isbased on the use of obstacle clearanceclimb speed.

e. Soft Field Takeoff. If a takeoff must bemade in conditions of mud, snow, tall grass, roughsurface or other conditions of high surface friction,the following procedure should be used. Set flaps atTAKEOFF (40%), align the aircraft with the runway,and with the yoke held firmly aft, begin a slowsteady acceleration, avoiding rapid or transientaccelerations. Continue to hold full aft yoke so as totransfer the weight of the aircraft from the wheels tothe wings as soon as possible. When the aircraftrotates, control pitch attitude (nose) so as to lift offfrom the soft surface at the slowest possible speed.When airborne, level off immediately in groundeffect just above the surface, and accelerate to nor-mal lift-off airspeed (V1of) before rotating to climbattitude and retracting the landing gear. Considerthe effects of snow or mud on gear retraction asapplicable.

8-42. AFTER TAKEOFF.

Immediately after takeoff, the pilot flyingthe aircraft should avoid adjusting con-trols located on the aft portion of theextended pedestal to preclude inducingspatial disorientation due to Coriolis illu-sion.

After the aircraft is positively airborne and flightis assured, retract the landing gear. Adjust pitch atti-tude as required to maintain best rate-of-climb air-speed (V,). Retract flaps after attaining best single-engine rate-of-climb airspeed (V,,). The copilotshould continue to maintain power at the computedsetting and to monitor instruments. At single-enginemaneuvering altitude, adjust pitch attitude to obtaincruise climb airspeed. As cruise climb airspeed, isattained, adjust power to the climb power setting.The copilot then activates the yaw damp and checksthat the cabin is pressurizing. Both pilots check thewings and nacelles for fuel or oil leaks. The proce-dural steps after takeoff are as follows:

1. Gear-up.

2. Flaps-UP.

3. Landing lights - OFF.

4. Climb power - Set.

5. PROP SYNC Switch - As required

06. YAW DAMP switch - As required.

0

Autofeather switch - As required.

Brake deice - As required.

9 Windshield Anti-ice - As required.

NOTE

Turn windshield anti-ice on to normal whenpassing 10,000 feet AGL or prior to entering thefreezing level (whichever comes first). Leaveon until no longer required during descent forlanding. High temperature may be selected asrequired, after a minimum warm-up period of15 minutes.

10. Cabin pressurization - Check, adjust RATE controlknob so that cabin rate-of-climb equals one-thirdaircraft rate-of-climb.

011. Wings and nacelles - Check.

012. Flare/chaff dispenser safety pin (electronicmodule) - Remove.

0Chaff function selector switch - As required.

14. APR-39 and APR-44 - As required.

8-43. CLIMB.

a. Cruise Climb. Cruise climb is performed at aspeed which is the best combination of climb, fuelbum-off, and distance covered Set propellers at1900 RPM and torque as required. Adhere to thefollowing airspeed schedule as closely as possible.

SL to 10,000 feet 140 KIAS10,000 to 20,000 feet 131 KIAS20,000 to 31,000 feet 121 KIAS

b. Climb - Maximum Rate. Maximum rate ofclimb performance is obtained by setting propellersat 2,000 RPM. torque at 100% (or maximum climbTGT), and maintaining best rate-of-climb airspeed.Refer to Chapter 7 for best rate-of-climb airspeedfor specific weights.

8-44. CRUISE.

Cruise power settings are entirely dependent uponthe 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.

TM 55-1510-221-10

1. Power - Set. Refer to the cruise power graphs con-tained in Chapter 7. To account for ram airtemperature increase, it is essential that tempera-ture be obtained at stabilized cruise airspeed.

NOTE

A new engine operated at the torque valuepresented in the cruise power charts will show aTGT margin below the maximum cruise limit.Maximum cruise power settings for temperatureand altitude (derived from Chapter 7) ifexceeded will adversely affect engine life. Withice vanes retracted, if cruise torque settingsshown on cruise power charts cannot be ob-tained without exceeding TGT limits, the en-gine should be inspected.

2. Ice & rain switches - As required. Insure that anti-ice equipment is activated before entering icingconditions.

NOTE

Ice vanes must be extended when operating invisible moisture at +5°C or less. Visible mois-ture is moisture in any form (clouds, ice crys-tals, snow, rain, sleet, hail, or any combinationof these).

03. Auxiliary fuel gages - Monitor. Insure that fuel isbeing transferred from auxiliary tanks. (Chapter 2,Section IV.)

04. Altimeters - Check. Verify that altimeter settingcomplies with transition altitude requirement.

05. Engine instrument indications - Check all engineinstruments for normal indications.

6. Recognition lights - As required.

8-45. DESCENT.

Descent from cruising altitude should normally bemade by letting down at cruise airspeed withreduced power. Refer to Chapter 7 for power set-tings and rates of descent.

NOTE

Cabin altitude and rat-f-climb controllershould be adjusted prior to starting descent.

a. Descent - Max Rate (Clean). To obtain themaximum rate of descent in clean configuration, per-form the following:

01. Cabin pressurization - Set. Adjust CABIN CON-TROLLER dial as required and adjust RATE con-trol knob so that cabin rate of descent equalsone-third aircraft rate of descent.

Change 3 8-19

TM 55-1510-221-10



4. Flaps - UP.

5. Gear - UP.

6. Airspeed - V mo.

07. Ice & rain switches - As requited.


b. Descent - Max Rate (Landing Configuration).If required to descend at a low airspeed (e.g., toconserve airspace or in turbulence), approach flapsand landing gear may be extended to increase therate and angle of descent while maintaining theslower airspeed. To perform the maximum rate ofdescent in landing configuration, perform the follow-ing:

01. Cabin pressurization - Set. Adjust CABIN CON-TROLLER dial as requited and adjust RATE con-trol knob so that cabin descent rate equalsone-third aircraft descent rate.



4. Flaps - APPROACH.

5. Gear - DN.

6. Airspeed - 180 KIAS maximum.

07. Ice & rain switches - As required.


8-46. DESCENT-ARRIVAL.

5. Altimeters - Set to current altimeter setting.

6. Flare/chaff dispenser arm-safe switch - SAFE.

7. Flare/chaff dispenser safety pin (electronicmodule) - Insert.

* 8. Crew briefing - Complete.

8-47. BEFORE LANDING.

1. Propeller synchronization switch - OFF.

02. Autofeather switch - ARM.

3. Propeller levers - As required.

NOTE

Perform the following checks prior to the finaldescent for landing.

Cabin pressurization - Set. Adjust CABINCONTROLLER dial as required.

Ice & rain switches - As required.

Windshield anti-ice - As required.

NOTE

Set windshield ANTI-ICE to normal or high asrequired well before descent into icing condi-tions or into warm moist air to aid in defogging.Turn off windshield anti-ice when descent iscompleted to lower altitudes and when heatingis no longer required. This will preclude pos-sible wind screen distortions.

4. Recognition lights - ON.

During approach, propellers should be set to1900 RPM to prevent glideslope interference(ILS approach), provide better power responseduring approach, and minimize attitude changewhen advancing propeller levers for landing.

4. Flaps (below 198 KIAS) - APPROACH.

5. Gear - DN.

6. Landing lights - As required.

07. Brake deice - As required.

8-48. OBSTACLE CLEARANCE APPROACH ANDMINIMUM RUN LANDING.

When landing over obstacles that require asteeper than normal approach path, or when greaterprecision is required due to restricted runwaylengths, the “Power Approach/Precision Landing”technique should be employed as follows: Prior tointercepting the descent path, complete the LAND-ING check and stabilize airspeed (V ref at 1.2 timespower-off stall speed in landing configuration (V so).

After intercepting the desired approach angle main-tain a constant descent by controlling the descentwith power and airspeed with elevator. Transitionsmoothly from approach to landing attitude. Touch-down should be made on the main gear with thenose slightly high, with power as required to controlrate of descent for a smooth touchdown. Immedi-ately after touchdown. allow the nosewheel to makeground contact and apply full reverse power andbraking, as required. If possible, remove reversethrust as the aircraft slows to 40 KIAS to minimizepropeller blade erosion.

NOTE

Using 1.2 x V so for approach airspeed will pro-vide increased performance and more respon-sive controls however, performance data are notavailable for approach at this slower airspeed.

8-20 Change 3

TM 55-1510-221-10

8-49. LANDING.

CAUTION

Except in an emergency, propellers should bemoved out of reverse above 40 knots tominimize propeller blade erosion, and duringcrosswind to minimize stress imposed onpropeller, engine and airframe. Care must beexercised when reversing on runways withloose sand and/or dust on the surface. Flyinggravel will damage propeller blades and dustmay impair the pilot’s forward visibility at lowairplane speeds. Performance data charts forlanding computations assume that the runwayis paved, level and dry. Additional runwaymust be allowed when these conditions are notmet. Refer to Chapter 7 for landing data. Donot consider headwind during landingcomputations; however, if landing must bedownwind, include the tailwind in landingdistance computations. Plan the final approachto arrive at 50 feet over the landing area atapproach speed (V ref) plus 1/2 wind gustspeed. Perform the following procedures as thean-craft nears the runway:

1.

2.

3.

a.

Autopilot and yaw damp - Disengaged.

Gear down lights - Check three green.

Propeller levers - HIGH RPM.

Normal. Landing. As the aircraft nears therunway, flare slightly to break the rate or descent andreduce power smoothly to idle as the nose. of the aircraftis rotated to landing attitude. Avoid the tendency to ridethe ground effect cushion while waiting for the aircraftto slow down to a soft landing. As the aircraft touchesdown, gently lower the nosewheel to the runway anduse reversing, brakes, or beta range, as required. Ifreversing is used, remove reverse power as the aircraftslows to 40 KIAS to minimize propeller blade erosion.

b. Crosswind Landing. Refer to Chapter 7 forrecommended Vref speeds. Use the "crab-into-thewind" method to correct for drift during final approach.The "crab" is changed to a slip (aileron into wind andtop rudder) to correct for drift during flare andtouchdown. After landing, position ailerons as requiredto correct for crosswind effect. For crosswind exceedingthe published limits, a combination "slip and crab"method at touchdown should be used.

c. Soft Field Landing. When landing on a softunprepared surface such as mud, tall grass, or snow,plan a normal power approach with flaps fully extended.Decelerate to the slowest possible airspeed just prior totouchdown, using power to control the final rate ofdescent to as slow as possible. Do not stall prior totouchdown as the nose attitude and rate of descent willbecome unacceptable. On touchdown apply full back(aft) elevator and then reduce power slowly. Do not use

brakes unless absolutely necessary. Every precautionmust be taken to prevent the nose wheel from digginginto the surface.

d. Touch-and-Go Landings. The instructor shouldselect a point on the runway where all pretakeoffprocedures will have been completed prior to the pilot’sinitial application of power. In selecting this point,prime consideration shall be given to the requiredaccelerate-stop distance pre-computed for the runway inuse. The nosewheel should be on the runway and rollingstraight before the touch-and-go procedures areinitiated. After the pilot applies power to within 5percent of target, the copilot’s (instructor’s) actions arethe same as during a normal takeoff. If touch-and-golandings are approved for training purposes use thefollowing procedure:

01 Propeller levers - HIGH RPM.

02 Flaps - As required.

03 Trim - Set.

4. Power stabilized - Check approximately 25%minimum.

5. Takeoff power - Set.

8-50. GO-AROUND.

When a go-around is commenced prior to theLANDING check, use power as required to climb to, ormaintain, the desired altitude and airspeed. If thego-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 theV ref speed used for the approach. Perform thefollowing:

1. Power - As required.

2. Gear - UP.

3. Flaps - UP.


5. Climb power - Set.

8-21

TM 55-1510-221-10

06. Yaw damp - As required.

Brake deice-OFF.

8-51. AFTER LANDING.

Complete the following procedures after theaircraft has cleared the runway:

02.

03

08.

CONDITION levers - As required.

Engine autoignition switch - OFF.

Ice & rain switches - OFF.

Flaps - UP.

Transponder - As required.

Radar - As required.

Lights - As required.

Mission control panel - Set.

8-52. ENGINE SHUTDOWN.

CAUTION

To prevent sustained loads on rudder shocklinks, the aircraft should be parked with thenose gear centered.

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

Brake deice - OFF.

Parking brake - Set.

Landing/taxi lights - OFF.

Cabin temperature mode selector switch -OFF.

Autofeather switch - OFF.

Vent and aft vent blower switches -AUTO.

INS - OFF.

Mission equipment - OFF, as required.

Inverter switches - OFF.

Battery condition - Check as required.

TGT - Check. TGT must be 660°C or belowfor one minute prior to shutdown.

CAUTION

Monitor TGT during shutdown. If sustainedcombustion is observed, proceed immediatelyto ABORT START procedure.

8-22 Change 4

12. Propeller levers - FEATHER.

13. CONDITION levers - FUEL CUTOFF.

WARNING

Do not turn exterior lights off until propeller’srotation has stopped.

14. Exterior lights - OFF.

15. Master panel lights switch - OFF.

16. Avionics master switch - Off.

17. Master switch - OFF. -

18. Keylock switch - OFF.

19. Oxygen system - OFF.

8-53. BEFORE LEAVING AIRCRAFT.

1. Wheels - Chocked.

2. Parking brake - As required.

NOTE

Brakes should be released after chocks are inplace (ramp conditions permitting).

3. Flight controls - Locked.

4. Overhead flood lights - Off.

5. Standby fuel pump switches - OFF.

6. Transponder - OFF.

7. Mode 4 - As required.

8. KY-28/58 - Zeroize as required.

9. Emergency exit lock - As required.

10. Aft cabin light - OFF.

11. Door light - OFF.

CAUTION

If strong winds are anticipated while theaircraft is unattended, the propellers shall besecured to prevent their windmilling with zeroengine oil pressure.

12. Walk-around inspection - Complete. Conducta thorough walk-around inspection. Checkingfor damage, fluid leaks. and levels. Check thatcovers, tiedowns. restraints.

TM 55-1510-221-10

safety pins and chocks are installed asrequired.

13.

NOTE

A cold oil check is unreliable. Oil shouldbe checked within 10 minutes after stopping engine. 14.

Aircraft forms - Complete. In addition toestablished requirements for reporting anysystem defects, unusual and excessive opera-tion such as hard landings, etc., the flightcrew will also make entries on DA Form2408-13 to indicate when limits in the Oper-ator’s Manual have been exceeded.

Aircraft - Secured. Lock cabin door asrequired.

Section III. INSTRUMENT FLIGHT

8-54. GENERAL.

This aircraft is qualified for operation in instru-ment meteorological conditions. Flight handling,stability characteristics and range are approximatelythe same during instrument flight conditions aswhen under visual flight conditions.

tion the flight director pitch steering bar. Retractflaps after attaining best single-engine rate-of-climbspeed (V yse), and re-adjust pitch as required. Controlbank attitude to maintain the desired heading. Sup-port flight director indications throughout themaneuver by crosschecking “raw data“ informationdisplayed on supporting instruments.

8-55. INSTRUMENT FLIGHT PROCEDURES. NOTE

Refer to FM 1-5, FM 1-230; FLIP; AR 95-1; FC1-2 18; or applicable foreign government regulations,and procedures described in this manual.

8-56. INSTRUMENT TAKEOFF.

Complete the BEFORE TAKEOFF check.Engage the heading (HDG) mode on the autopilotcomputer/control (do not engage autopilot). Setheading marker (HDG) to runway heading and alignthe aircraft with the runway centerline, insuring thatnosewheel is straight before stopping aircraft. Holdbrakes and complete the LINEUP check. Insure thatthe roll steering bar is centered. Power applicationand copilot duties are identical to those prescribedfor a “visual” takeoff. After the brakes are released,initial directional control should be accomplishedpredominantly with the aid of outside visual refer-ences. As the takeoff progresses, the crosscheckshould transition from outside references to theflight director and airspeed indicator. The rate oftransition is directly proportional to the rate atwhich the outside references deteriorate. Approach-ing rotation speed (V,), the crosscheck should betotally committed to the instruments so that errone-ous sensory inputs can be ignored. At rotationspeed, establish takeoff attitude on the flight direc-tor. Maintain this pitch attitude and wings-level atti-tude until the aircraft becomes airborne. When boththe vertical-velocity indicator and altimeter showpositive climb indications, retract the landing gear.After the landing gear is retracted, adjust the pitchattitude as required to attain best rate-of-climb air-speed (V,). Use PITCH-SYNC as required to reposi-

Due to possible precession error, the pitchsteering bar may lower slightly duringacceleration, causing the pitch attitude toappear higher than actual pitch attitude.To avoid lowering the nose prematurely,crosscheck the vertical-velocity indicatorand altimeter to insure proper climb per-formance. The erection system will auto-matically remove the error after the accel-eration ceases.

8-57. INSTRUMENT CLIMB.

Instrument climb procedures are the same asthose for visual climb. Enroute instrument climbsare normally performed at cruise climb airspeeds.

8-58. INSTRUMENT CRUISE.

There are no unusual flight characteristics dur-ing cruise in instrument meteorological conditions.

8-59. INSTRUMENT DESCENT.

When a descent at slower than recommendedspeed is desired, slow the aircraft to the desiredspeed before initiating the descent. Normal descentto approach altitude can be made using cruise air-speed. Normally, descent will be made with the air-craft in a cruise configuration, maintaining desiredspeed as required.

8-23

TM 55-1510-221-10

8-60. INSTRUMENT APPROACHES.

There are no unusual preparations or controltechniques required for instrument approaches. Theapproaches are normally flown at an airspeed of V ref+20 until transitioning to visual flight.

8-61. AUTOPILOT APPROACHES.

There are no special preparations required forplacing the aircraft under autopilot control. Refer toChapter 3 for procedures to be followed for auto-matic approaches.

NOTE

The ILS localizer and glideslope warningflags indicate insufficient signal strengthto the receiver. Certain electrical mechan-ical malfunctions between the receiverand indicators may result in erroneouslocalizer/glideslope information without awarning flag. It is recommended that ILSinformation be crosschecked with otherflight instruments prior to and duringfinal approach. Utilization of NAV TESTon VOR control prior to the f inalapproach fix may detect certain malfunc-tions not indicated by the warning flags.

Section III. FLIGHT CHARACTERISTICS

8-62. STALLS.

A prestall warning in the form of very light buf-feting can be felt when a stall is approached. Anaural warning is provided by a warning horn. Thewarning horn starts sounding approximately five toten knots above stall speed with the aircraft in anyconfiguration. If correct stall recovery technique isused, very little altitude will be lost during the stallrecovery. For the purpose of this section, the term"power-on" shall mean that both engines and pro-pellers of the aircraft are operating normally andpower is set at approximately 50%. The term"power-off" shall mean that both engines are operat-ing at idle power. Landing gear position has noeffect on stall speed. During practice, enter power-off stalls from normal glides. Enter power-on stallsby smoothly increasing pitch attitude to decreaseairspeed by approximately one knot per second untilstall occurs.

a. Power-On Stalls. The power-on stall atti-tude is very steep and unless this high-pitch attitudeis maintained, the aircraft will generally “settle“ or“mush“ instead of stall. It is difficult to stall the air-craft inadvertently in any normal maneuver. A lightbuffet precedes most stalls, and the first indicationof approaching stall is generally a decrease in controleffectiveness, accompanied by a “chirping“ tonefrom the stall warning horn. The stall itself is char-acterized by a rolling tendency to the right, if theaircraft is allowed to yaw. The proper use of rudderwill prevent the tendency to roll. A slight pitchingtendency will develop if the aircraft is held in thestall, resulting in the nose dropping sharply, thenpitching up toward the horizon; this cycle isrepeated until recovery is made. Control is regained

very quickly with little altitude loss, providing thenose is not lowered excessively. Begin recovery withforward movement of the control wheel and a grad-ual return to level flight. The roll tendency causedby yaw is more pronounced in power-on stalls, as isthe pitching tendency; however, both are easily con-trolled after the initial entry. Power-on stall charac-teristics are not greatly affected by wing flap posi-tion, except that stalling speed is reduced inproportion to the degree of wing flap extension.

b. Power-Off Stalls. Power-off stalls are char-acterized by a right rolling tendency, as the stall isapproached. Elevator control is effective to the stopand the pitch attitude can be maintained with adeceleration rate of 1 knot/sec. Light to moderatebuffet commences approximately 5 - 8 knots abovethe stall and the warning horn will sound and con-tinue to the stall. With wing flaps down, the rightrolling tendency is more pronounced and stallingspeed is much slower than with the wing flaps up.The Stall Speed Chart (fig. 8-2) shows the indicatedpower-off stall speeds with the aircraft in variousconfigurations. Altitude loss during a full stall willbe approximately 800 feet.

c. Accelerated Stalls. The aircraft gives notice-able stall warning in the form of buffeting when thestall occurs. The stall warning and buffet can bedemonstrated in turns by applying excessive backpressure on the control wheel.

8-63. SPINS.

Intentional spins are prohibited. If a spin isinadvertently entered use the following recoveryprocedure:

8-24

Figure 8-2. Stall Speed

8-25

TM 55-1510-221-10

NOTE 8-65. MANEUVERING FLIGHT.

Spin demonstrations have not been con-ducted. The recovery technique is basedon the best available information. Thefirst three actions should be accomplishedas nearly simultaneous as possible.

Maneuvering speed (V,), 168 KIAS), is the max-imum speed that abrupt control movements can beapplied without exceeding the design load factor ofthe aircraft. The data is based on 15,000 pounds.


2. Apply full rudder opposite the direction ofspin rotation.

8-66. FLIGHT CONTROLS.

3. Simultaneously with rudder application,push the control wheel forward and neutral-ize ailerons.

4. When rotation stops, neutralize rudder.

Do not pull out of the resulting dive tooabruptly as this could cause excessivewing loads and 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.

8-64. DIVING.

Maximum diving airspeed (red line) is 243KIAS or 0.47 Mach. Flight characteristics are con-ventional 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.

The aircraft is stable under all normal flight con-ditions. Aileron, elevator, rudder and trim tab con-trols function effectively throughout all normal flightconditions. Elevator control forces are relativelylight in the extreme aft CG (center of gravity) condi-tion, progressing to moderately high with CG at theforward limit. Extending and retracting the landinggear causes only slight changes in control pressure.Control pressures, resulting from changes in powersettings or the repositioning of the wing flaps are notexcessive in the landing configuration at the mostforward CG position. The minimum speed at whichthe aircraft can be fully trimmed is 89 KIAS (gearand flaps down, propellers at high RPM, and 15,000pounds power for a 3° angle of descent. Controlforces produced by changes in speed, power setting,wing flap position and landing gear position are lightand can be overcome with one hand on the controlwheel. Trim tabs permit the pilot to reduce theseforces to zero. During single engine operation, therudder boost system aids in relieving the relativelyhigh rudder pressures resulting from the large varia-tion in power.

8-67. LEVEL FLIGHT CHARACTERISTICS.

All flight characteristics are conventionalthroughout the level flight speed range.

Section V. ADVERSE ENVIRONMENTAL CONDITIONS

8-68. INTRODUCTION.

The purpose of this part is to inform the pilot ofthe special precautions and procedures to be fol-lowed during the various weather conditions that

may 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.

8-26

TM 55-1510-221-10

8-69. COLD WEATHER OPERATIONS.

Operation of the surface deice system inambient temperatures below -40°C cancause permanent damage to the deiceboots. Operational diffIculties may beencountered during extremely coldweather, unless proper steps are takenprior to or immediately after flight. Allpersonnel should understand and be fullyaware of the necessary procedures andprecautions 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, anti-ice solution, or brake deice, to free frozen tires.When heat is applied to release tires, the tempera-ture should not exceed 71 °C (160°F). Refer to Chap-ter 2 for anti-icing, deicing, and defrosting treat-ment.

b. Engine Starting. When starting engines onramps covered with ice, propeller levers should be inthe FEATHER position to prevent the tires fromsliding. To prevent exceeding torque limits whenadvancing CONDITION levers to HIGH IDLE dur-ing the starting procedure, place the power lever inBETA and the propeller lever in HIGH RPM beforeadvancing the condition lever to HI IDLE.

c. Warm-Up and Ground Test. Warm-up pro-cedures and ground test are the same as those out-lined in Section II.

d. Taxiing. Whenever possible, taxiing indeep snow, light weight dry snow or slush should beavoided, particularly in colder OAT conditions. If itis necessary 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 rolling

while parked. Before attempting to taxi, activate thebrake deice system, insuring that the bleed air valvesare OPEN and that the condition levers are in HIIDLE. 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.

(1.) If icing conditions are expected, acti-vate all anti-ice systems before takeoff, allowing suf-ficient time for the equipment to become effective.

(2.) If the possibility of ice accumulationon the horizontal stabilizer or elevator exists, takeoffwill not be attempted. If icing conditions areexpected, activate all anti-ice systems before takeoff,allowing sufficient time for the equipment tobecome effective.

f. Takeoff. Takeoff procedures for coldweather operations are the same as for normal take-off. Taking off with temperatures at or below freez-ing, with water, slush or snow on the runway, cancause ice to accumulate on the landing gear and canthrow ice into the wheel well areas. Such takeoffsshall be made with brake deice on and with the icevanes extended to preclude the possibility of icebuild-up on engine air inlets. Monitor oil tempera-tures to insure operation within limits. Before flightinto icing conditions, the pilot and copilot WSHLDANTI-ICE switches should be set at NORMAL posi-tion.

g. During Flight.

(1.) Brake deice. After takeoff from a run-way covered with snow or slush, it may be advisableto leave brake deice ON to dislodge ice accumulatedfrom the spray of slush or water. Monitor BRAKEDE-ICE annunciator for automatic termination ofsystem operation and then turn the switch OFF.

(2.) Flight controls. During flight, trimtabs and controls should also be exercised periodi-cally to prevent freezing.

(3.) Anti-icing equipment. Insure that anti-icing systems are activated before entering icing con-ditions. 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.

8-27

TM 55-1510-221-10

NOTE

Do not operate deicer boots continuously.Continuous operation tends to balloonthe ice over the boots. Allow at least 1/2inch of ice to accumulate on the surfaceboots and 1/8 to 1/4 inch of ice to accu-mulate on the antenna boots, then acti-vate the deicer boots to remove the ice.Repeat this procedure as required.

(4.) Ice vanes. Ice vanes must be extendedwhen operating in visible moisture or when freedomfrom visible moisture cannot be assured, at 5°COAT or less. Ice vanes are designed as an anti-icesystem, not a deice system. After the engine air inletscreens are blocked, lowering the ice vanes will notrectify the condition. Ice vanes should be retractedat 15°C OAT and above to assure adequate engineoil cooling.

(5.) Stall speeds. Stalling airspeeds shouldbe expected to increase when ice has accumulatedon the aircraft causing distortion of the wing airfoil.For the same reason, stall warning devices are notaccurate and should not be relied upon. Keep a com-fortable margin of airspeed above the normal stallairspeed. Maintain a minimum of 140 KIAS duringsustained icing conditions to prevent ice accumula-tion on unprotected surfaces of the wing. In theevent of windshield icing, reduce airspeed to 226KIAS or below.

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 within10 degrees of runway heading. Application of brakeswithout skidding the tires on ice is very difficult,due to the sensitive brakes. In order not to impairpilot visibility, reverse thrust should be used withcaution when landing on a runway covered withsnow or standing water. Use the procedures in Sec-tion II for normal landing.

j. Engine Shutdown. Use normal proceduresin Section II.

k. Before Leaving the Aircraft. When the air-craft 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, removeany accumulation of dirt and ice from the landing

8-28

gear shock struts, and install protective covers toguard against possible collection of snow and ice.

8-70. DESERT OPERATION AND HOT WEATHEROPERATION.

Dust, sand, and high temperatures encounteredduring desert operation can sharply reduce the oper-ational life of the aircraft and its equipment. Theabrasive qualities of dust and sand upon turbineblades and moving parts of the aircraft and thedestructive effect of heat upon the aircraft instru-ments will necessitate hours of maintenance if basicpreventive measures are not followed. In flight, thehazards of dust and sand will be difficult to escape,since dust clouds over a desert may be found at alti-tudes up to 10,000 feet. During hot weather opera-tions, the principle difficulties encountered are highturbine gas temperatures (TGT) during engine start-ing, over-heating of brakes, and longer takeoff andlanding rolls due to the higher density altitudes. Inareas where high humidity is encountered, electricalequipment (such as communication equipment andinstruments) will be subject to malfunction by corro-sion, fungi and moisture absorption by nonmetallicmaterials.

a. Preparation For Flight. Check the positionof the aircraft in relation to other aircraft. Propellersand blast can damage closely parked aircraft. Checkthat the landing gear shock struts are free of dustand sand. Check instrument panel and general inte-rior for dust and sand accumulation. Open mainentrance door and cockpit vent storm windows toventilate the aircraft.

N1 speeds of 70% or higher may berequired to keep oil temperature withinlimits.

b. Engine Starting. Use normal procedures inSection II. Engine starting under conditions of highambient temperatures may produce a higher thannormal TGT during the start. The TGT should beclosely monitored when the condition lever ismoved to the LO IDLE position. If overtemperaturetendencies are encountered, the condition levershould be moved to the IDLE CUTOFF positionperiodically during acceleration of gas generatorRPM (N1). Be prepared to abort the start beforetemperature limitations are exceeded.

c. Warm-Up Ground Tests. Use normal proce-dures in Section II. To minimize the possibility ofdamage to the engines during dusty/sandy condi-

TM 55-1510-221 -10

tlons. activate ICE VANES if the temperature is below15°C.

d. Taxing. Use normal procedures in Section II.When practical. avoid taxiing over sandy terrain tominimize propeller damage and engine deterioration thatresults from Impingement of sand and gravel. During hotweather operation. use minimum braking action to preventoverheating.

e. Takeoff Use normal procedures in Section II.Avoid taking off in the wake of another alrcraft of therunway surface is sandy or dusty.

II. f. During Flight. Use normal procedures in Section

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 expansion 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 brake immediately after installingwheel chocks to prevent brake disc warpage.

8-71. 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 allnecessary changes in aircraft attitude with theleast amount of control pressures to avoidexcessive loads on the aircraft’s structure.

Thunderstorms and areas of severe turbulence should beavoided. If such areas are to be penetrated. it will benecessary to counter rapid changes in attitude and acceptmajor Indicated altitude variations. Penetration should beof an altitude which provides adequate maneuveringmargins as a loss or gain of several thousand feet of

altitude may be expected. The recommended penetrationspeed in severe turbulence is 158 KIAS. Pitch attitude andconstant power settings are vital to proper flight technique.Establish recommended penetration speed and properattitude prior to entering turbulent air to minimize mostdifficulties. False Indications by the pressure Instrumentsdue to barometric pressure variations within the stormmake them unreliable. Maintaining a pre-establishedattitude will result in a fairly constant airspeed. Turncockpit and cabin lights on to minimize the blinding effectsof lighting. Do not use autopilot altitude hold. Maintalnconstant power settings and pitch attitude regardless ofairspeed or altitude indications. Concentrate on maintaininga level attitude by reference to the Flight Director/AttitudeIndicator. Maintain original heading. Maker no turnsunless absolutely necessary.

8-72. ICE AND RAIN (TYPICAL).

WARNING

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 recommended 1/2-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 Ievel flight. This can be determined by comparing pre-

Change 5 8-28.1

TM 55-1510-221-10

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 for it to stick.If severe icing conditions are encountered, ascend ordescend to altitudes where these conditions do not prevail.If flight into icing conditions is unavoidable, proper use ofaircraft anti-icing and deicing systems may minimize theproblems encountered. Approximately 15 minutes prior toflight into temperature conditions which could producefrost or icing conditions, the pilot and co-pilot windshieldanti-ice switches should be set at normal or hightemperature position (after preheating) as necessary toeliminate windshield ice. Stalling airspeeds should beexpected to increase when ice has accumulated on theaircraft causing distortion of the wing airfoil. For the samereason. stall warning devices are not accurate and shouldnot be relied upon. Keep a comfortable margin of airspeedabove the normal stall airspeed with ice on the aircraft.Maintain a minimum of 140 knots during sustained icingconditions to prevent ice accumulation on unprotectedsurfaces of the wing. In the event of windshield icing,reduce airspeed to 226 knots or below.

b. Rain. Rain presents no particular problems otherthan restricted visibility and occasional incorrect airspeedindications.

c. Taxiing. Extreme care 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. Cruise Flight.

(1) Prevention of ice formation. Prevention of iceformation is far more effective and satisfactory thanattempts to dislodge the ice after it has formed. If icing

conditions are inadvertently encountered, turn on the anti-icing systems prior to the first sign of ice formation.

(2) Deicer boots. Do not operate deicer bootscontinuously. Allow at least one-half inch of ice on theboots before activating the deicer boots to remove the ice.Continued flight in severe icing conditions should not beattempted. If ice forms on the wing area aft of the deicerboots, climb or descend to an altitude where conditions areless 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.

NOTE

When operating on wet or icy runways, refer tostopping distance factors shown in Chapter 7.

8-72A. 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 5-30B.), 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 avoid

8-28.2 Change 5

TM 55-1510-221-10

extended exposure to flight conditions more severe thanthose for which the airplane has been certificated.

(b) Avoid abrupt and excessive maneuveringthat may exacerbate control difficulties

(f) Do not cxtend 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.

(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.

(g) If the flaps are extended. do not retractthem until the airframe is clear of ice.

(h) Report these weather conditions to airtraffic control.

Section VI. CREW DUTIES

* 8-73. CREW BRIEFING.

The following guide should be used in accomplishinprequired 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.

2. Emergency equipment

3. Emergency landing/ditching procedures.

* 8-74. DEPARTURE BRIEFING.

The following is a guide that should be used asapplicable in accomplishing the required crew briefingprior to takeoff. However. if the crew has operatedtogether previously and the pilot is certain that the copilotunderstands all items of the briefing. he may omit thebriefing by stating “standard briefing.” when the briefing iscalled for during the BEFORE TAKEOFF CHECK.

a. ATC clearance - Review.

1. Routing

Change 5 8-29/(8-30 blank)

2. Initial altitude.

b. Departure procedure - Review.

1. SID.

2. Noise abatement procedure.

3. VFR departure route.

c. Copilot duties - Review.

1.

2.

3.

4.

5.

6.

7.

8.

9.

Adjust takeoff power.

Monitor engine instruments.

Power check at 65 KIAS.

Call out engine malfunctions.

Tune/ident all nav/com radios.

Make all radio calls.

Adjust transponder and radar asrequired.

Complete flight log during flight (notealtitudes and headings).

Note departure time.

d. PPC - Review.

1. Takeoff power.

2 . V r

3. V y, (climb to 500’ AGL).

4 . V y s e

8-75. ARRIVAL BRIEFING.

The following is a guide that should be used asapplicable in accomplishing the required crew brief-ing prior to landing; however, if the crew has oper-ated together previously and the pilot is certain that

TM 55-1510-221-10

the copilot understands all items of the briefing, hemay omit the briefing by stating “standard briefing,“when the briefing is called for during theDESCENT-ARRIVAL check.

a. Weather/altimeter setting.

b. Airfield/facilities - Review.

2. 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. Call out visual/field in sight.

f . Landing performance data - Review.

1. Approach speed.

2. Runway required.

8-31/(8-32 blank)

TM 55-1510-221-10

CHAPTER 9

EMERGENCY PROCEDURES

Section I. AIRCRAFT SYSTEMS

9-1. AIRCRAFT SYSTEMS. 9-4. AFTER EMERGENCY ACTION.

This section describes the aircraft systems emer-gencies that may reasonably be expected to occurand presents the procedures to be followed. Emer-gency procedures are given in checklist form whenapplicable. A condensed version of these proceduresis in the Operator’s and Crewmember’s Checklist,TM 55-1510-221-CL. Emergency operations of avi-onics equipment are covered when appropriate inChapter 3, Avionics, and are repeated in this sectiononly as safety of flight is affected.

After a malfunction has occurred, appropriateemergency actions have been taken, and the aircraftis on the ground, an entry shall be made in theremarks section of DA Form 2408-13 describing themalfunction.

9-5. EMERGENCY EXITS AND EQUIPMENT.

Emergency exits and equipment are shown infigure 9-1.

9-6. EMERGENCY ENTRANCE.9-2. IMMEDIATE ACTION EMERGENCYCHECKS.

Immediate action emergency items are under-lined for your reference and shall be committed tomemory. During an emergency, the checklist will becalled for to verify the memory steps performed andto assist in completing any additional emergencyprocedures.

Entry may be made through the cabin emer-gency hatch. The hatch may be released by pullingon its flush-mounted pull-out handle, placardedEMERGENCY 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.

NOTE 9-7. ENGINE MALFUNCTION.

The urgency of certain emergenciesrequires immediate action by the pilot.The most important single considerationis aircraft control. All procedures are sub-ordinate to this requirement. Reset MAS-TER CAUTION after each malfunctionto allow systems to respond to subsequentmalfunctions.

9-3. DEFINITION OF LANDING TERMS.

The term LANDING IMMEDIATELY isdefined as executing a landing without delay. (Theprimary consideration is to assure the survival ofoccupants.) The term LAND AS SOON AS POSSI-BLE is defined as executing a landing at the nearestsuitable landing area without delay. The termLAND AS SOON AS PRACTICABLE is defined asexecuting a landing to the nearest suitable airfield.

a. Flight Characteristics Under Partial PowerConditions. There are no unusual flight characteris-tics during single-engine operation as long as air-speed is maintained at or above minimum controlspeed (V,,) and above power-off stall speed. Thecapability of the aircraft to climb or maintain levelflight depends on configuration, gross weight, alti-tude, and outside air temperature. Performance andcontrol will improve by feathering the propeller ofthe inoperative engine, retracting the landing gearand flaps, and establishing the appropriate single-engine best rate-of-climb speed (V yse). Minimumcontrol speed (V,,) with flaps retracted is approxi-mately 1 knot 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 (V 1of) has been attained.If an engine fails immediately after liftoff, manyvariables such as airspeed, runway remaining, air-

9-1

TM 55-1510-221-10

Figure 9-1. Emergency Exits and Equipment

9 - 2

craft 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). If anengine fails and the aircraft has not accelerated torecommended liftoff speed (V 1of), retard power leversimmediately to IDLE and stop the aircraft with brakesand reverse thrust. Perform the following:

1. Power levers - IDLE.

2. Braking - As required.

NOTE

If able to land on remaining runway, checkgear down and use brakes and reversethrust as required. If insufficient runwayremains for stopping, perform the following:

3. Condition levers - FUEL CUTOFF.

4. Fire pull handles - Pull.

5. Master switch - OFF.

d. Engine Malfunction After Liftoff. If an enginefails after becoming airborne, maintain single-enginebest rate-of-climb speed (V yse) or, if air speed is belowV yse maintain whatever airspeed is attained betweenliftoff (V 1of) and V yse

until sufficient altitude is attainedto trade altitude for-airspeed and accelerate to V yse.

1.

2.

3.

(1.) Engine Malfunction after liftoff (abort),perform the following and land in awingslevel attitude:

Power levers - Reduce.

Gear - DN.

Complete normal landing.

NOTE

If able to land on remaining runway, checkgear down and use brakes and reversethrust as required. If insufficient runwayremains for stopping, perform the following:

4. Condition levers - FUEL CUTOFF



(2.) Engine malfunction after liftoff (flightcontinued) perform the following:

TM 55-1510-221-10

1. Power - Maximum controllable.

NOTE

If airspeed is below V yse maintain whateverairspeed has been attained (between V 1of andV yse) until sufficient altitude can be obtainedto trade off altitude for airspeed to assist inacceleration to V yse.

2. Gear - UP

3. Flaps - UP.


5. Brake deice - OFF.

6. Engine cleanup - Perform.

7. Generator load - 100% max.

NOTE

Holding three to five degrees bank (one-halfball width) towards the operating engine willassist in maintaining directional control andimprove aircraft performance.

e. Engine Malfunction During Flight. If an enginemalfunctions during cruise flight, maintain control of theaircraft while maintaining heading or turn as required.Add power as required to keep airspeed from decayingexcessively and to maintain altitude. Identify the failedengine by feel (if holding rudder pressure to keep theaircraft from yawing; the rudder being pressed indicatesthe good engine) and engine instruments, then confirmidentification by retarding the power lever of thesuspected failed engine. Refer to Chapter 7 forsingle-engine cruise information. If one enginemalfunctions during flight, perform the following:

1. Autopilot/yaw damp - DISENGAGE.


3. Dead engine - Identified.

4. Power lever (dead engine) - IDLE.

5. Propeller lever (dead engine) - FEATHER.

6. Propeller synchronization switch - Off.

7. Gear - As required.


9. Generator load - 100% max.

10. Power - Set for single engine cruise.

Change 4 9 -3

TM 55-1510-221-10

11. Engine cleanup - Perform.

NOTE

At V yse, speeds, holding three to five degreesbank (one-half ball width) towards theoperating engine will assist’ in maintainindirectional control and improve aircraftperformance.

f. Engine Malfunction During Final Approach. Ifan engine malfunctions during finalLANDING CHECK) the propeller

approach (aftershould not be

manually feathered unless time and altitude permit orconditions require it. Continue approach using thefollowing procedure:

1. Power-As required

2. Gear - DN.

g. Engine Malfunction (Second Engine). If thesecond engine fails, do not feather the propeller if anengine restart is to be attempted. Engine restart withoutstarter assist can not be accomp lished with a featheredpropeller, and the propeller Will not unfeather withoutthe engine operatingbest all around glid

140 KIAS is recommended as thee speed (considering engine restart,

distance covered, transition to landing configuration,etc.), although it does not necessarily result in theminimum rate of descent. Perform the followingprocedure if the second engine fails during cruise flight.

1. Airspeed - 140 KIAS.

2. Powerlever - IDLE

3. Propeller lever - Do not FEATHER.

4. Conduct engine restart procedure.

h. Engine Shutdown In Flight. If it becomesnecessary to shut an engine down during flight, performthe following:

1. Power lever - IDLE.

2. Propeller lever - FEATHER.

3. Condition lever - FUEL CUTOFF.

4. Engine clean up - Perform.

i. Engine Cleanup. The cleanused after engine malfunction,unsuccessful restart is as follows:

1. Autoignition switch - OFF.

2. Autofeather switch - OFF.

3. Generator switch - OFF.

up procedure to beshutdown, or an

4. Propeller synchronization switch - OFF.

9-4 Change 4

j. Engine Restart During Flight Using Starter.Engine restarts may be attempted at all altitudes. Ifrestart is attempted, perform the following:

1.

2.

3.

4.

5.

6.

7.

8.

9.

Cabin temperature mode selector switch - OFF.

Electrical load - Reduce to minimum.

Fire pull handle - In.

Power lever - IDLE.

Propeller lever - FEATHER.

Condition lever - FUEL CUTOFF.

TGT (operative engine) - 700°C or less.

Ignition and engine start switch - ON.

Condition lever - LOW IDLE.

NOTE

If a rise in TGT does not occur within 10seconds after moving the condition lever toLOW IDLE, abort the start.

10. TGT - Monitor (1,000°C for 5 secondsmaximum).

11. Oil pressure - Check.

12. Ignition and engine start switch - OFF at 50%N1.

13. Generator switch - RESET, then ON.

14. Engine cleanup - Perform if engine restartunsuccessful.

15. Cabin temperature mode selector switch - Asrequired.

16. Electrical equipment - As required.

17. Autoignition switch - ARM.

18. Propellers - Synchronize.


k. Engine Restart During Flight (Not UsingStarter). A restart without starter assist may beaccomplished provided airspeed is at or above 140KIAS altitude is below 20,000 feet, and the propeller isnot feathered. If altitude permits, diving the aircraft willincrease N1 and assist in restart. N1 required for airstartshould be at or above 9%. If a start is attempted,perform following:

1. Cabin temperature mode selector switch - OFF.

2.

3.

Electrical load - Reduce to minimum.

Generator switch (affected engine) -OFF.

4. Fire pull handle - Check in.


6. Propeller lever - HIGH RPM.

7. Condition lever - FUEL CUTOFF.

8. Airspeed - 140 KIAS minimum.

9. Altitude below 20,000 feet - Check.

10. Engine autoignition switch - ARM.

11. Condition lever - LOW IDLE.

NOTE

If N, 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.

NOTE

If a rise in TGT does not occur within 10seconds after moving the condition leverto LOW IDLE, abort the start.

12.

13.

14.

15.

16.

17.

18.

19.

20.

TGT - Monitor (1 ,OOO”C for 5 secondsmaximum).

Oil pressure - Check.

Generator switch - RESET, then ON.

Engine Cleanup - Perform if enginerestart unsuccessful.

Cabin temperature mode selectorswitch - As required.

Electrical equipment - As required.

Autoignition switch - ARM.

Propellers - Synchronized.

Power - As required.

l. Maximum Glide. In the event of failure ofboth engines, maximum gliding distance can beobtained by feathering both propellers to reducepropeller drag and by maintaining the appropriateairspeed with the gear and flaps up. Figure 9-2 givesthe approximate gliding distances in relation to alti-tude.

Maizainm. Landing With Two Engines Inoperative.

best glide speed (figure 9-2). If sufficient

TM 55-1510-221-10

altitude remains after reaching a suitable landingarea, a circular pattern will provide best observationof surface conditions, wind velocity, and direction.When the condition of the terrain has been notedand the landing area selected, set up a rectangularpattern. Extending APPROACH flaps and landinggear early in the pattern will give an indication ofglide performance sooner and will allow more timeto make adjustments for the added drag. Fly thebase leg as necessary to control point of touchdown.Plan to overshoot rather than undershoot, then useflaps as necessary to arrive at the selected landingpoint. Keep in mind that, with both propellers feath-ered the normal tendency is to overshoot due to lessdrag. In event a positive gear-down indication can-not be determined, prepare for a gear-up landing;also, unless the surface of the landing area is hardand smooth, the landing should be made with thelanding gear up. If landing on rough terrain, land ina slightly tail-low attitude to keep nacelles from pos-sibly digging in. If possible, land with flaps fullyextended.

9-8. LOW OIL PRESSURE.

In the event of a low oil pressure indication, per-form the procedures below as applicable:

1. Oil pressure below 105 PSI below 21,000feet or 85 PSI 21,000 feet and above, torque- 49% maximum.

2. Oil pressure below 60 PSI - Perform engineshutdown, or land as soon as practicableusing minimum power to insure safe arrival.

9-9. CHIP DETECTOR WARNING LIGHT ILLUMI-NATED.

If a L CHIP DETR or a R CHIP DETR warninglight illuminates, and safe single-engine flight can bemaintained; perform engine shutdown.

9-10. DUCT OVERTEMP CAUTION ANNUNCIA-TOR LIGHT ILLUMINATED.

If a DUCT OVERTEMP caution annunciatorlight is illuminated, insure that the cabin floor out-lets are open and unobstructed, then perform thefollowing steps in sequence until the light is extin-guished. After completion of steps 1 thru 4, if lightdoes not extinguish, Allow approximately 30 sec-onds after each adjustment for the system tempera-ture to stabilize. The overtemperature condition isconsidered corrected at any point during the proce-dure that the light extinguishes.

1. Cabin air control - In.

9-5

TM 55-1510-221-10

Figure 9-2. Maximum Glide Distance

9 - 6

2.

3.

4.

5.

6.

7.

8.

9.

10.

Cabin temperature mode selector switch -AUTO.

Cabin temperature control rheostat - Fulldecrease.

Vent blower switch - HI.

Cabin temperature mode selector switch -MAN COOL.

Manual temperature switch - DECREASE(hold).

Left bleed air valve switch - ENVIRO OFF.

If the light is still illuminated after 30 sec-onds: Left bleed air valve switch - OPEN.

Right bleed air valve switch - ENVIROOFF.

If the light is still illuminated after 30 sec-onds: Right bleed air valve switch - OPEN.

NOTE

If the overtemperature light has not extin-guished after completing the above proce-dure, the warning system has malfunc-tioned.

9-11. ICE VANE FAILURE.

Ice vane failure is indicated by #1 VANE FAILor #2 VANE FAIL caution annunciator light illumi-nation. If an ice vane fails to operate electrically,perform the following:

After the ice vanes have been manuallyextended, they must be mechanicallyretracted. No electrical extension orretraction shall be attempted as damageto the electric actuator may result. Link-age in the nacelle area must be reset priorto operation of the electric system. Do notreset ice vane control circuit breaker.

Do not retract ice vanes electrically aftermanual extension.

1. Airspeed - 160 KIAS or below.

2. Ice vane control circuit breaker - Pull.

TM 55-1510-221-10

3. Ice vane - Operate manually.

4. Airspeed - Resume normal airspeed.

9-12. ENGINE BLEED AIR SYSTEM FAILURE.

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 theremainder of flight. Perform the following:

NOTE

L BL AIR FAIL or R BL AIR FAIL lightsmay momentarily illuminate duringsimultaneous surface deice and brakedeice operation at low N1 speeds.


2. TGT and torque - Monitor (note read-ings).

3. Bleed air valve switch - PNEU &ENVIRO OFF.

NOTE

Brake deice on the affected side, and rud-der boost, will not be available with bleedair valve switch in PNEU & ENVIROOFF.

4. Cabin pressurization - Check.

b. Excessive Differential Pressure. If cabin dif-ferential pressure exceeds 6.1 PSI, perform the fol-lowing:

1. Cabin altitude and rate-of-climb con-troller -Select higher setting.

If condition persists: LEFT BLEEDAIR VALVE switch - ENVIRO OFF(light illuminated).

2.

3.

4.

5.

6.

If condition still persists: RIGHTBLEED AIR VALVE switch -ENVIRO OFF (light illuminated).

If condition still persists - Descendimmediately.

If unable to descend: CABIN PRESSDUMP switch -CABIN PRESSDUMP.

Bleed air valve switches - OPEN, ifcabin heating is required.

9-7

TM 55-1510-221-10

9-13. LOSS OF PRESSURIZATION (ABOVE 10,000 FEET).

If cabin pressurization is lost when operatingabove 10,000 feet or the ALT WARN warningannunciator light illuminates, perform the following:

1. Crew oxygen masks - 100% and on.

9-14. CABIN DOOR CAUTION LIGHT ILLUMI-NATED.

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.

9-15. SINGLE-ENGINE DESCENT/ARRIVAL.

NOTE

Approximately 85% N1 is required tomaintain pressurization schedule.

Perform the following procedure prior to thefinal descent for landing.

1. Cabin controller - Set.

2. Ice and rain switches - As required.

3. Altimeters - Set.

4. Recognition lights - ON.

* 5. Arrival briefing - Complete.

9-16. SINGLE-ENGINE BEFORE LANDING.

1. Propeller lever - As required.

NOTE

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.


3. Gear - DN.

4. Landing lights - As required.

9-8

5. Yaw damp - OFF.


9-17. SINGLE-ENGINE LANDING CHECK.

Perform the following procedure during finalapproach to runway.

1. Autopilot/yaw damp - Disengaged.

2. Gear lights - Check (three green).

3. Propeller lever (operative engine) - HIGHRPM.

NOTE

To insure constant reversing characteris-tics, the propeller control must be in theHIGH RPM position.

9-18. SINGLE-ENGINE GO-AROUND.

The decision to go around must be made asearly as possible. Elevator forces at the start of a go-around are very high and a considerable amount ofrudder control will also be required at low airspeeds.Retrim as required. If rudder application is insuffi-cient, or applied too slowly, directional control can-not be maintained. If control difficulties are experi-enced, reduce power on the operating engineimmediately. Insure that the aircraft does not touchthe ground before retracting the landing gear.Retract the flaps only as safe airspeed permits(TAKEOFF until V ref, then UP).engine go-around as follows:

Perform single-

Once flaps are fully extended, a single-engine go-around may not be possiblewhen close to the ground under condi-tions of high gross weights and/or highdensity altitude.

1. Power - Maximum allowable.

2. Gear - UP.




6. Yaw damp - As required.

9-19. PROPELLER FAILURE (OVER 2080 RPM).

If an overspeed condition occurs that cannot becontrolled with the propeller lever, or by reducingpower, perform the following:

1. Power lever (affected engine) - IDLE.

2. Propeller lever - FEATHER.

3. Condition lever - As required.

4. Propeller synchronization - OFF.

5. Engine cleanup - As required.

9-20. FIRE.

The safety of aircraft occupants is the primaryconsideration when a tire occurs; therefore, it isimperative that every effort be made by the flightcrew to put the fire out. On the ground it is essentialthat the engines be shut down, crew evacuated, andfire fighting 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/nacelle fire during start orground operations. If engine/nacelle fire is identifiedduring start or ground operation, perform the fol-lowing:

1. Propeller levers - FEATHER.

2. Condition levers - FUEL CUT-OFF.

3. Fire pull handle - Pull.

If fire extinguisher has been used to extin-guish an engine fire, do not attempt torestart, until maintenance personnel haveinspected the aircraft and released it forflight.

4. Push to extinguish switch - Push.


(2.) Engine fire in flight fire pull handlelight illuminated). If an engine fire is suspected inflight, perform the following:


2. If fire pull handle light out isextinguished: Advance power.

TM 55-1510-221-10

3. If fire pull handle light is still illu-minated: Engine fire in flight pro-cedures (identified) - Perform.

NOTE

Flight into the sun at high aircraft pitchattitude may actuate the tire warning sys-tem. Lowering the nose and/or changingheadings will confirm a warning systemfailure-caused by sun rays.

(3.) Engine fire in frightengine fire is confirmed in flight,lowing:

(identified). If anperform the fol-

Due to the possibilities of fire warningsystem malfunctions, the fire should bevisually identified before the engine issecured and the extinguisher actuated.

1.

2.

3.

4.

5.

6.

Power lever - IDLE.

Propeller lever - FEATHER.

Condition lever - FUEL CUT-OFF.

Fire pull handle - Pull.

Fire extinguisher - Actuate asrequired.

Engine cleanup - Perform.

b. Fuselage Fire. If a fuselage fire occurs, per-form the following:

The extinguisher agent (Bromochlorodi-fluoromethane) in the fire extinguishercan produce toxic effects if inhaled.

1. Fight the fire.

2. Land as soon as possible.

c. Wing Fire. There is little that can be doneto control a wing fire except to shut off fuel andelectrical systems that may be contributing to thefire, or which could aggravate it. Diving and slippingthe aircraft away from the burning wing may help.If a wing fire occurs, perform the following:

1. Perform engine shutdown on affectedside.

9-9

7. Essential electrical equipment - On (in-dividually until fire source is isolated).

8. Land as soon as practicable.

e. Smoke and Fume Elimination. To elimi-nate smoke and fumes from the aircraft, perform thefollowing:

1.

2.

Crew oxygen - 100% and ON.

Bleed air valve switches - ENVIROOFF.

3.

4.

5.

Vent blower switch - AUTO.

Aft vent blower switch - OFF.

Cabin temperature mode selectorswitch - OFF.

6. If smoke and fumes are not eliminated:Cabin pressure dump switch - CABINPRESS DUMP.

TM 55-1510-221-10


d. Electrical Fire. Upon noting the existenceor indications of an electrical fire, turn off allaffected electrical circuits, if known. If electrical firesource is unknown, perform the following:

1. Crew oxygen - 100%.

2. Master switch - OFF (visual conditionsonly).

3. All nonessential electrical equipment -OFF.

NOTE

With loss of DC electrical power, the air-craft will depressurize. All electricalinstruments, with the exception of thepropeller RPM, Nt RPM, and TGT gageswill be inoperative.

4. Battery switch - ON.

5. Generator switches (individually) -RESET, then ON.

6. Circuit breakers - Check for indicationof defective circuit.

As each electrical switch is returned toON (note loadmeter reading) and checkfor evidence of fire.

NOTE

Opening storm window (after depressuriz-ing) will facilitate smoke and fumeremoval.

7. Engine oil pressure - Monitor.

9-21. FUEL SYSTEM.

a. Fuel Pressure Warning Annunciator LightIlluminated. Illumination of the #1 FUEL PRESS or#2 FUEL PRESS warning light usually indicates fail-ure of the respective engine-driven boost pump. Per-form the following:

1. Standby pump switch - ON.

2. Fuel pressure warning annunciatorlight - Check extinguished.

3. If fuel pressure warning light is stillilluminated: Record unboosted time.

b. No Fuel Transfer Caution Light Illumi-nated. Illumination of a #1 NO FUEL XFR or #2NO FUEL XFR annunciator light with fuel remain-ing in the respective auxiliary fuel tank indicates afailure of that automatic fuel transfer system. Pro-ceed as follows:

1. AUX TRANSFER switch (affectedside) - OVERRIDE.

2. Auxiliary fuel quantity - Monitor.

3. AUX TRANSFER switch (after respec-tive auxiliary fuel has completely trans-ferred) - AUTO.

c. Nacelle Fuel Leak. If nacelle fuel leaks areevident, perform the following:

1. Perform engine shutdown.



d. Fuel Crossfeed. Fuel crossfeed is normallyused only during single-engine operation. The fuelfrom the dead engine side may be used to supply thelive engine by routing the fuel through the crossfeedsystem. During extended flights, this method of fuelusage will provide a more balanced lateral load con-dition in the aircraft. For fuel crossfeed, use the fol-lowing procedure:

1. AUX TRANSFER switches - AUTO.

9-10

NOTE

With the FIRE PULL handle pulled. thefuel in the auxiliary tank for that side willnot be available (usable) for crossfeed.

2. Standby pumps - OFF.

3. Crossfeed switch - As required.

4. Fuel crossfeed advisory annunciatorlight - Check illuminated.

NOTE

With the FIRE PULL handle pulled, therespective WI FUEL PRESS or #2 FUELPRESS light will remain illuminated onthe side supplying fuel.

5. Fuel pressure light extinguished -Check.

6. Fuel quantity - Monitor.

e. Illumination of the #1 NAC LOW or #2NAC LOW caution annunciator fight. Illuminationof the #1 NAC LOW or #2 NAC LOW cautionannunciator light indicates that the affected tank has20 minutes remaining at sea level, normal cruisepower consumption rate. Proceed as follows:

Failure of the fuel tank venting system willprevent the fuel in the wing tanks from gravityfeeding into the nacell 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 Iittle as 114 pounds, regardless of thetotaI fuel quantity indicated. Continued fIightmay result in engine flameout due to fuel star-vation.

1. Twenty minutes fuel remaining - Con-firm:


NOTE

If a "NAC LOW" light occurs about thetime the "AUX" tanks go empty and thefuel gages show the main tanks "FULL".

TM 55-1510-221-10

this may indicate the fuel vent float valvein the wet section has stuck. Rocking theaircraft or changing pitch will probablyunstick it. If not. fuel may be crossfed.

9-22 ELECTRICAL SYSTEM EMERGENCIES.

a. DC Generator Caution Annunciator LightIlluminated. Illumination of a #1 DC GEN or #2DC GEN caution annunciator light indicates failureof a generator or one of its associated circuits (gener-ator control unit). If one generator system becomesinoperative. all nonessential electrical equipmentshould be used judiciously to avoid overloading theremaining generator. The use of accessories whichcreate a very high drain should be avoided. If bothgenerators are shut off due to either generator sys-tem failure or engine failure. all nonessential equipment should be turned off to preserve battery powerfor extending the landing gear and wing flaps. Whena DC GEN light illuminates, perform the following:

1. Generator switch - OFF, RESET, thenON.

2. Generator switch (no reset) - OFF.

3. Mission control switch - OVERRIDE.

4. Operating loadmeter - 100% maxi-mum.

b. Both DC Generator Warning AnnunciatorLights Illuminated.

1. All nonessential equipment - OFF.


c. Excessive Loadmeter Indication (Over100%). If either loadmeter indicates over 100%, per-form the following:

1. Battery switch - OFF (monitor load-meter).

2. Loadmeter over 100% - Nonessentialelectrical equipment OFF.

3. Loadmeter under 100% - BATT switchON.

d. Inventer Caution Annunciator Light Illumi-nated. Illumination of the #1 INVERTER or #2lNVERTER caution annunciator light indicates fail-ure of the affected inverter. When either inverterfails, the total aircraft load is automatically switchedto the remaining inverter. When a #1 INVERTERor #2 INVERTER caution annunciator light illumi-nates, perform the following:

1. A f fec ted #1 INVERTER or #2INVERTER switch - OFF.

Change 1 9-11

TM 55-1510-221-10

e. INST AC Warning Annunciator Light Illu-minated. Illumination of the INST AC warning lightindicates that 26 VAC power is not available. Allitems connected to the 26 VAC bus will be inopera-tive (refer to AC wiring schematic diagram in chapter 2 for equipment effected). Under these condi-tions. power must be controlled by indications of theNt and TGT gages. Perform the following

1. Nt and TGT indications - Check.

2. Other engine instruments - Monitor.

f. Circuit Breaker Tripped. If the circuitbreaker is for a nonessential item. do not reset inflight. If the circuit breaker is for an essential item,the circuit breaker may be reset once. If a bus feedercircuit breaker (on the overhead circuit breakerpanel) trips. a short is indicated. Do not reset inflight. If a circuit breaker trips. perform as follows:

1. BUS FEEDER breaker tripped - Donot reset.

2. Nonessential circuit - Do not reset.

3-. Essential circuit - Reset once.

NOTE

Circuit breakers should not be reset morethan once until the cause of the circuitmalfunction has been determined andcorrected. Do not reset dual fed busfeeder circuit breakers.

g. BATTERY CHARGE Light Illuminated. Ifthe BATTERY CHARGE caution light illuminatesduring normal cruise flight. perform the following:

1. Battery Volt-Ampmeter - Monitor. Ifbattery current continues to increase,turn battery switch off.

2. Battery switch (landing gear/flap exten-sion only) - ON.

9-23. EMERGENCY DESCENT.

Emergency descent is a maximum effort inwhich damage to the aircraft must be consideredsecondary to getting the aircraft down. The follow-ing procedure assumes the structural integrity of theaircraft and smooth flight conditions. If structuralintegrity is in doubt. limit speed as much as possi-ble. reduce rate of descent if necessary, and avoidhigh maneuvering loads. For emergency descent,perform the following:

NOTE

Windshield defogging may be required.


2. Propeller lever - HIGH RPM.


4. Gear - DN.

5. Airspeed - 180 KIAS maximum.

9-24. LANDING EMERGENCIES.

Structural damage may exist after landingwith brake, tire, or landing gear malfunc-tions. Under no circumstances shall anattempt be made to inspect the aircraftuntil jacks have been installed.

a. Landing Gear Unsafe Indication. Shouldone or more of the three green landing gear indica-tor lights fail to indicate a safe condition. the follow-ing steps should be taken before proceeding toextend the gear manually.

1. Gear - DN.


3. Landing gear relay and indicator cir-cuit breaker - Check In.

NOTE

If gear continues to indicate unsafe.attempt to verify position of the landinggear visually.

h. Landing Gear Emergency Extension.

Continued pumping of the handle afterGEAR DOWN position indicator lights(3) are illuminated could damage thedrive mechanism. and prevent subsequentgear retraction.

9-12 Change 1

TM 55-1510-221-10

After an emergency landing gear exten-sion has been made. do not stow the gearratchet handle or move any landing gearcontrols or reset any switches or circuitbreakers until the cause of the malfunc-tion has been corrected.


2. LANDING GEAR RELAY circuitbreaker - Out.

3. Gear - DN.

4. Landing gear alternate engage handle -Lift and turn clockwise to the stop.

5. Alternate landing gear extension handle- Pump.


c. Gear-up Landing (All Gear Up orUnlocked). Due to decreased drag with the gear up.the tendency will be to overshoot the approach. The

Change 1 9-12.1/(9-12.2 blank)

center-of-gravity with the gear retracted is aft of themain wheels. This condition will allow the aircraftto be landed with the gear retracted and shouldresult in a minimum amount of structural damage tothe aircraft, 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 isnot recommended as sod has a tendency to roll upinto chunks, damaging the underside of the aircraft’sstructure. When fuel load has been reduced, preparefor a gear-up landing as follows:

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

Crew emergency briefing - Complete.

Loose equipment - Stowed.

Bleed air valve switches - ENVIROOFF.

Cabin pressure dump switch - CABINPRESS DUMP.

Cabin emergency hatch - Remove andstow.

Seat belts and harnesses - Secured.

Landing gear alternate engage handle -Disengaged.

Alternate landing gear extension handle- Stowed.

Gear relay circuit breaker - In.

Gear - UP.

Nonessential electrical equipment -OFF.

Flaps - As required (DOWN for land-ing).

NOTE

Fly a normal approach to touchdown.After landing, accomplish the following:

13. Power levers (runway assured) - IDLE.




d. Landing With Nose Gear Unsafe. If thelanding gear control switch handle warning light isilluminated and the nose GEAR DOWN indicatorlight shows an unsafe condition, the nose gear isprobably not locked down, and the gear positionshould be checked visually by another aircraft, ifpossible. If all attempts to lock the nose gear fail, alanding should be made with the main gear downand locked. Hold the nose off the runway as long as

TM 55-1510-221-10

possible and do not use brakes. Use the followingprocedures:

1. Crew emergency briefing - Complete.

2. Loose equipment - Stowed.

3. Bleed air valve switches - ENVIROOFF.

4. Cabin pressure dump switch - CABINPRESS DUMP.

5. Cabin emergency hatch - Remove andstow.

6. Seat belts and harnesses - Secured.

7. Nonessential electrical equipment -OFF.

NOTE






e. Landing With One Main Gear Unsafe. Ifone main landing gear fails to extend, retract theother gear and make a gear-up landing. If all effortsto retract the extended gear fail, land the aircraft ona hard runway surface, touching down on the sameedge of the runway as the extended gear. Roll on thedown and locked gear, holding the opposite wing upand the nose gear straight as long as possible. If thegear has extended, but is unsafe, apply brakes lightlyon the unsafe side to assist in locking the gear. If thegear has not extended or does not lock, allow thewing to lower slowly to the runway. Use the follow-ing procedures:

1. Crew emergency briefing - Complete.

2. Loose equipment - Stowed.

3. Bleed air valve switches - ENVIROOFF.

4. Cabin pressure dump switch - CABINPRESS DUMP.

5. Cabin emergency hatch - Remove andstow.

6. Seat belts and harnesses - Secured.

7. Nonessential electrical equipment -OFF.

9-13

TM 55-1510-221-10

8. Touchdown - On safe main gear first.

NOTE






f. Landing With Flat Tire(s). If aware that amain gear tire(s) is flat, a landing close to the edgeof the runway opposite the flat tire will help avoidveering off the runway. If the nose wheel tire is flat,use minimum braking.

9-25. LANDING WITH INOPERATIVE WINGFLAPS (UP).

The aircraft does not exhibit any unusual char-acteristics when landing with the wing flaps up. Theapproach angle will be shallow and the touchdownspeed will be higher resulting in a longer landingroll.

9-26. CRACKED WINDSHIELD.

a. External Crack. If an external windshieldcrack is noted, no action is required in flight.

NOTE

Heating elements may be inoperative inareas of crack.

b. Internal Crack. If an internal crack occurs,perform the following:

1. Descend to below 25,000 feet.

2. Cabin Pressure - Reset pressure differ-ential to 4 PSI or less within 10 min-utes.

9-27. CRACKED CABIN WINDOW (OUTERPANEL).

If a cabin window outer panel crack occurs, per-form the following:

1. Descend to below 25,000 feet.

2. Cabin pressure - 4.6 PSI maximum.

3. Do not operate more than 20 flight hours.

NOTE

Treat outer panel cracks which are linear(not circular) or cracks that touch theframe, as an inner panel crack.

9-28. CRACKED CABIN WINDOW (INNERPANEL).

If a cabin window inner panel crack occurs, per-form the following:

1. Oxygen - As required.

2. Cabin pressure - Depressurize.

3. Descend - As required.

9-29. DITCHING.

If a decision to ditch is made, immediately alertall crewmembers to prepare for ditching. Plan theapproach 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 mini-mum rate descent (power on or off, as the situationdictates, airspeed - (110-120 KIAS). Do not try toflare as in a normal landing, as it is very difficult tojudge altitude over water, particularly in a slick sea.Leveling off too high may cause a nose low “dropin,“ while having the tail too low on impact mayresult in the aircraft pitching forward and “diggingin.“ Expect more than one impact shock and severalskips before the final hard shock. There may benothing but spray visible for several seconds whilethe aircraft is decelerating. To prevent cartwheeling,it is important that the wings be level when the air-craft hits the water. After the aircraft is at rest,supervise evacuation of passengers and exit the air-craft as quickly as possible. In a planned ditching,the life raft and first-aid kits should be secured closeto the cabin 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 damaged sur-faces which might tear or puncture the fabric. Thelength of time that the aircraft will float depends onthe fuel level and the extent of aircraft damagecaused by the ditching. Refer to figure 9-3 for bodypositions during ditching. Figure 9-4 shows windswell information. Perform the following proce-dures:

9-14

Do not unstrap from the seat until allmotion stops. The possibility of injuryand disorientation requires that evacua-tion not be attempted until the aircraftcomes to a complete stop.

1.

2.

3.

4.

Radio calls/transponder - As required.

Crew emergency briefing - As required.

Bleed air valve switches - ENVIRO OFF.

Cabin pressure dump switch - CABINPRESS DUMP.

The rudder boost system may not operatewhen the brake deice system is in use.Availability of the rudder boost systemwill be restored to normal when theBRAKE DEICE switch is turned off.

5.

6.

7.

8.

9.

10.

11.

Cabin emergency hatch - Remove and stow.

Seat belts and harnesses - Secured.

Gear - UP.

Flaps - DOWN.

Nonessential electrical equipment - OFF.

Approach - Normal, power on.

Emergency lights - As required.

IF CONDITION PERSISTS:

2. Bleed air valve switches - PNEU &ENVIRO OFF.

3. Rudder trim - Adjust.

c. Unscheduled Electric Elevator Trim. In theevent of unscheduled electric elevator trim, performthe following:

1. Elevator trim switch - OFF.

2. Elevator trim circuit breaker - Out.

9-30 FLIGHT CONTROLS MALFUNCTION. 9-31. BAILOUT.

Use the following procedures, as applicable, forflight control malfunctions.

a. Autopilot/Yaw Damp Emergency Discon-nection. The autopilot can be disengaged by any ofthe following methods:

I.

2.

3.

4.

5.

6.

Pressing the DISC - TRIM - AP - YDdisconnect switch (control wheels).

Pressing the autopilot "AP ENGAGE"pushbutton on the autopilot modeselector control panel.

Pressing the go-around switch (leftpower lever), (yaw damper will remainon).

Pulling the AP CONTR and AFCSDIRECT circuit breakers (overheadcontrol panel).

Setting AVIONICS MASTER PWRswitch (overhead control panel) to theOFF position.

Setting aircraft MASTER switch (over-head control panel) to the OFF posi-tion.

TM 55-1510-221-10

b. Unscheduled Rudder Boost Activation. Rud-der boost operation without a large variation ofpower between engines indicates a failure of the sys-tem. Perform the following:

1. Rudder boost - OFF.

NOTE

When the decision has been made to abandonthe aircraft in flight, the pilot will give the warningsignal. Exit from the aircraft will be through themain entrance door, and in the departure sequenceusing the exit routes as indicated in figure 9-1. Pro-ceed as follows if bailout becomes necessary:

1. Notify crew to prepare to bail out.

2. Distress message - Transmit.

3. Voice security - ZEROIZE.

4. Transponder - 7700.

5. Mode 4 - Zeroize.

6. Flaps - DOWN.


8. Trim - As required.

9. Autopilot - Engage.

10. Cabin pressure switch - DUMP.

11. Parachute - Attach to harness.

12. Cabin door - Open.

13. Abandon the aircraft.

9-15

TM 55-1510-221-10

Table 9-1. Ditching

PLANNED DITCHINGPILOT

A. ALERT OCCUPANTSB. ORDER TO PREPARE SURVIVAL GEAR FOR

AERIAL DROPC. TRANSMIT DISTRESS MESSAGED. LIFE VEST - CHECK (DO NOT INFLATE)E. DISCHARGE MARKER

F. LAND AND DITCH AIRCRAFTG. ABANDON AIRCRAFT

IMMEDIATE DITCHINGPILOT

A. WARN OCCUPANTSB. TRANSMIT DISTRESS MESSAGE

C. LIFE VEST - CHECK (DO NOT INFLATE)D. APPROACH - NORMALE. NOTIFY OCCUPANTS TO BRACE FOR

DITCHINGF. LAND AND DITCH AIRCRAFTG. ABANDON AIRCRAFT AFTER COPILOT

THROUGH CABIN EMERGENCY HATCH

COPILOT COPILOT

A. REMOVE CABIN EMERGENCY HATCH A. REMOVE CABIN EMERGENCY HATCHB. LIFE VEST - CHECK (DO NOT INFLATE) B. LIFE VEST - CHECK (DO NOT INFLATE)C. ABANDON AIRCRAFT (TAKE LIFE RAFT AND C. ABANDON AIRCRAFT (TAKE LIFE RAFT AND

FIRST AID KIT) FIRST AID KIT)

PASSENGERS PASSENGERS

A. SEAT BELTS - FASTEN A. SEAT BELTS - FASTENB. LIFE VEST - CHECK ( DO NOT INFLATE B. LIFE VEST - CHECK (DO NOT INFLATE)C. ON PILOTS SIGNAL - BRACE FOR DITCHING C. ON PILOTS SIGNAL -BRACE FOR DITCHINGD. ABANDON AIRCRAFT THROUGH CABIN D. ABANDON AIRCRAFT THROUGH CABIN

DOOR (TAKE LIFE RAFT AND FIRST AID KIT DOOR (TAKE LIFE RAFT AND FIRST AID KIT)

9-16

TM 55-1510-221-10

Figure 9-3. Emergency Body Positions

9-17

TM 55-1510-221-10

Figure 9-4. Wind Swell Ditch Heading Evaluation

9-18

TM 55-1510-221-10

APPENDIX A

REFERENCES

Reference information for the subject material contained in this manual can be found in the followingpublications:

AR 70-50

AR 95-1

AR 95-16

AR 380-40

AR 385-40

AR 700-26

DA PAM 738-751

FAR Part 91

FM 1-5

FM 1-30

TB AVN 23-13

TB MED 501

TB 55-9150-200-24

TM 9-1095-206-13&P

(C) TM 11-5825-252-15

TM 11-5841-291-12

TM 11-5841-283-20

TM 11-6140-203-14-2

TM 11-6940-214-12

TM 55-410

TM 55-1500-204-25/1

TM 55-1500-314-25

TM 55-1500-342-23

TM 55-1510-200-PM

TM 55-1510-219-23

Designating and Naming Defense Equipment, Rockets, and GuidedMissles

Army Aviation - General Provisions and Flight Regulations

Weight and Balance - Army Aircraft

Safeguarding COMSEC Information

Accident Reporting and Records

Aircraft Designation System

Functional User’s Manual for the Army Maintenance Management Sys-tem Aviation (TAMS-A)

General Operating and Flight Rules

Instrument Flying and Navigation for Army Aviators

Meterology for Army Aviators

Anti-icing, Deicing and Defrosting Procedures for Parked Aircraft

Noise and Conservation of Hearing

Engine and Transmission Oils, Fuels, and additives for Army Aircraft

Operator’s Aviation Unit Maintenance and Aviation IntermediateMaintenance Manual (Including Repair Parts and Special Tools List)to Dispenser, General Purpose Aircraft: M-130

Operator, Organizational, DS, GS, and Deport Maintenance Manual:RC-12H Aircraft Mission Equipment, (V)

Operator and Organizational Maintenance Manual, Radar WarningSystem, AN/APR-44(V)1

Organizational Maintenance Manual for Detection Set, Radar SignalAN/APR-39(V)1.

Operator’s Organizational, Direct Support, General Support and DepotMaintenance Manual Including Repair Parts and Special Tools List:Aircraft Nickel-Cadmium Batteries

Operator and Organization Maintenance Manual, Simulator, Radar Sig-nal, SM-756/APR-44(V)

Aircraft Maintenance, Servicing and Ground Handling Under ExtremeEnvironmental Conditions

General Aircraft Maintenance Manual Maintenance Manual: ArmyModel RC-12H Aircraft

Handling, Storage, and Disposal of Army Aircraft Components Con-taining Radioactive Materials

Army Aviation Maintenance Manual: Weight and Balance

Phased Maintenance Checklist

Aviation Unit and Aviation Intermediate

A-1

TM 55-1510-221-10

TM 750-244-1-5 Procedures for the Destruction of Aircraft and Associated Equipmentto Prevent Enemy Use

A-2

TM 55-1510-221-10

APPENDIX B

ABBREVIATIONS AND TERMS

For the purpose of this manual, the following abbreviations and terms apply. See appropriate technicalmanuals for additional terms and abbreviations.

Airspeed Terminology.

CAS

FT/MIN

GS

IAS

Calibrated airspeed is indicated airspeed corrected for position andinstrument error.

Feet per minute.

Ground speed, though not an airspeed, is directly calculable from trueairspeed if the true wind speed and direction are known.

Indicated airspeed is the speed as shown on the airspeed indicator andassumes no error.

KT

TAS

V a

V f

V fe

Knots.

True airspeed is calibrated airspeed corrected for temperature, pres-sure, and compressability effects.

Maneuvering speed is the maximum speed at which application of fullavailable aerodynamic control will not overstress the aircraft.

Design flap speed is the highest speed permissible at which wing flapsmay be actuated.

Maximum flap extended speed is the highest speed permissible withwing flaps in a prescribed extended position.

V le

V 1o

V lof

Vmea

VmoVne

vr

vs

Maximum landing gear extended speed is the maximum speed atwhich an aircraft can be safely flown with the landing gear extended.

Maximum landing gear operating speed is the maximum speed atwhich the landing gear can be safely extended or retracted.

Lift off speed (takeoff airspeed).

The minimum flight speed at which the aircraft is directionally con-trollable as determined in accordance with Federal Aviation Regula-tions. Aircraft Certification conditions include one engine becominginoperative and windmilling; a 5° bank towards the operative engine;takeoff power on operative engine; landing gear up; flaps up; and mostrearward CG. This speed has been demonstrated to provide satisfac-tory control above power off stall speed (which varies with weight,configuration, and flight attitude).

Maximum operating limit speed.

Never exceed speed.

Rotation speed.

V s o

Vsse

Power off stalling speed or the minimum steady flight speed at whichthe aircraft is controllable.

Stalling speed or the minimum steady flight speed in the landing con-figuration.

The safe one-engine inoperative speed selected to provide a reasonablemargin against the occurence of an unintentional stall when makingintentional engine cuts.

B-1

TM 55-1510-221-10

Best angle of climb speed.

Best single-engine angle of climb speed.

Best rate of climb speed.

The best single engine rate of climb speed.

Meteorological Terminology.

Altimeter Setting

°C

°F

FAT

Indicated Pressure Altitude

ISA International Standard Atmosphere in which:

Pressure Altitude

SL

Wind

Beta Range

Cruise Climb

High Idle

HP

Low Idle

Maximum Cruise Power

Maximum Power

Normal Rated Climb Power

Normal Rated Power

Reverse Thrust

B-2

Barometric pressure corrected to sea level.

Degrees Celsius.

Degrees Fahrenheit.

Free Air Temperature is the free air static temperature obtained eitherfrom the temperature indicator (IFAT), adjusted for compressibilityeffects, or from ground meteorlogical sources.

The number actually read from an altimeter when, the barometric scale(Kollsman window) has been set to 29.92 inches of mercury (1013 mil-libars).

a. The air is a dry perfect gas.

b. The temperature at sea level is 59 degrees Fahrenheit, 15degrees Celsius.

c. The pressure at sea level is 29.92 inches Hg.

d. The temperature gradient from sea level to the altitude at whichthe temperature is -69.7 degrees Fahrenheit is -0.003566 Fahren-heit per foot and zero above that altitude.

Indicated pressure altitude corrected for altimeter error.

Sea level.

The wind velocities recorded as variables on the charts of this manualare to be understood as the headwind or tailwind components of theactual winds at 50 feet above runway surface (tower winds).

The region of the power lever control which is aft of the idle stop andforward of reversing range where blade pitch angle can be changedwithout a change of gas generator RPM.

Is the maximum power approved for normal climb. This power istorque or temperature (ITT) limited.

Obtained by placing the condition lever in the HIGH IDLE position.

Horsepower.

Obtained by placing the condition lever in the LO IDLE position.

Is the highest power rating for cruise and is not time limited.

The maximum power available from an engine for use during an emer-gency operation.

The maximum power available from an engine for continuous normalclimb operations.

The maximum power available from an engine for continuous opera-tion in cruise (with lower ITT limit than normal rated climb power).

Obtained by lifting the power levers and moving them aft of the betarange.

TM 55-1510-221-10

RPM

Takeoff Power

Revolutions Per Minute.

The maximum power available from an engine for takeoff, limited toperiods of five minutes duration.

Control and Instrument Terminology.

Condition Lever (Fuel Shut-off The fuel shut-off lever actuates a valve in the fuel control unit whichLever) controls the flow of fuel at the fuel control outlet and regulates the idle

range from LO to HIGH.

Interstage Turbine Temperature Eight probes wired in parallel indicate the temperature between the(ITT) compressor and power turbines.

Nt Tachometer (Gas Generator The tachometer registers the RPM of the gas generator with 100% rep-RPM) resenting a gas generator speed of 37,500 RPM.

Power Lever (Gas Generator N1 This lever serves to modulate engine power from full reverse thrust toRPM) takeoff. The position for idle represents the lowest recommended level

of power for flight operation.

Propeller Control Lever (N2 This lever requests the control to maintain RPM at a selected valueRPM) and, in the maximum decrease RPM position, feathers the propeller.

Propeller Governor This Governor will maintain the selected propeller speed requested bythe propeller control lever.

Torquemeter The torquemeter system determines the shaft output torque. Torquevalues are obtained by tapping into two outlets on the reduction gearcase and recording the differential pressure from the outlets.

Graph and Tabular Terminology.

AGL

Best Angle of Climb

Best Rate of Climb

Clean Configuration

Demonstrated Crosswind

Gradient

Landing Weight

Maximum Zero Fuel Weight

MEA

Obstacle Clearance ClimbSpeed

Ramp Weight

Route Segment

Above ground level.

The best angle-of-climb speed is the airspeed which delivers the great-est gain of altitude in the shortest possible horizontal distance withgear and flaps up.

The best rate-of-climb speed is the airspeed which delivers the greatestgain of altitude in the shortest possible time with gear and flaps up.

Gear and flaps up regardless of mission antenna installation.

The maximum 90° crosswind component for which adequte control ofthe aircraft during takeoff and landing was actually demonstrated dur-ing certification tests.

The ratio of the change in height to the horizontal distance, usuallyexpressed in percent.

The weight of the aircraft at landing touchdown.

Any weight above the value given must be loaded as fuel.

Minimum Enroute Altitude.

Obstacle clearance climb speed is a speed near Vx and Vy 1.1 timespower off stall speed, or 1.2 times minimum single-engine stall-speed,whichever is higher.

The gross weight of the aircraft before engine start. Included is thetakeoff weight plus a fuel allowance for start, taxi, run up and takeoffgrond roll to liftoff.

A part of a route. Each end of that part is identified by:

a. A geographic location; or

B-3

TM 55-1510-221-10

b. 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 min-ute can be attained for existing aircraft weight.

Takeoff Weight The weight of the aircraft at liftoff from the runway.

Weight and Balance Terminology.

Arm

Approved Loading Envelope

Basic Empty Weight

Center-of-Gravity

CG Limits

Datum

Engine Oil

Empty Weight

Landing Weight

Maximum Weight

Moment

Standard

Station

Takeoff Weight

Unusable Fuel

Usable Fuel

Useful Load

Miscellaneous Abbreviations.

DegDN

FT

FT LB

The distance from the center of gravity of an object to a line aboutwhich moments are to be computed.

Those combinations of aircraft weight and center of gravity whichdefine the limits beyond which loading is not approved.

The aircraft weight with unusable fuel, full oil, and full operating flu-ids.

A point at which the weight of an object may be considered concen-trated for weight and balance purposes.

CG limits are the extremes of movement which the CG can have with-out making the aircraft unsafe to fly. The CG of the loaded aircraftmust be within these limits at takeoff, in the air, and on landing.

A vertical plane perpendicular to the aircraft longitudinal axis fromwhich fore and aft (usually aft) measurements are made for weight andbalance purposes.

That portion of the engine oil which can be drained from the engine.

The aircraft weight with fixed ballast, unusable fuel, engine oil, enginecoolant, hydraulic fluid, and in other respects as required by applicableregulatory standards.

The weight of the aircraft at landing touchdown.

The largest weight allowed by design, structural, performance or otherlimitations.

A measure of the rotational tendency of a weight, about a specifiedline, mathematically equal to the product of the weight and the arm.

Weights corresponding to the aircraft as offered with seating and inte-rior, avionics, accessories, fixed ballast and other equipment specifiedby the manufacturer as composing a standard aircraft.

The longitudinal distance from some point to the zero datum or zerofuselage station.

The weight of the aircraft at liftoff.

The fuel remaining after consumption of usable fuel.

That portion of the total fuel which is available for consumption asdetermined in accordance with applicable regulatory standards.

The difference between the aircraft ramp weight and basic emptyweight.

Degrees

Down

Foot or feet

Foot-pounds

B - 4

TM 55-1510-221-10

GAL Gallons

HR Hours

kHz Kilohertz

LB Pounds

MAX Maximum

MHz Megahertz

MIN Minimum

NAUT Nautical

NM Natucial miles

PSI Pounds per square inch

R/C Rate of climb

B-5/(B-6 blank)

TM 55-1510-221-10

INDEX

Paragraph, Figure,Table NumberSubject

A

Abort Start ...................................................... 8-32AC Power Supply ............................................ 2-75Accelerate-Go Distance Over 50-Foot

Obstacle (Flaps 0%) .................................... 7-10Accelerate-Stop (Flaps 0%) .............................. 7-7Accelerometer .................................................. 2-85Action Codes and Recommended

Actions .......................................................... T3-5Additional Data (Normal

Procedures) .................................................... 8-8After Emergency Action .................................... 9-4After Landing .................................................. 8-51After Takeoff .................................................... 8-42Air Conditioning System ................................ 2-70Air Induction Systems - General .................... 2-18Aircraft Compartment and

Stations ................................................ 6-3, F6-1Aircraft Dimensions ........................................ F6-1Aircraft Designation System .......................... 1-11Aircraft Systems ................................................ 9-1Airspeed Indicators .......................................... 2-81Airspeed Limitations ........................................ 5-19Altitude Limitations .......................................... 5-28Altitude Select Controller ................................ 3-23Ammunition ........................................................ 4-5Antenna Deicing System ................................ 2-52Anti-Icing, Deicing and Defrosting

Treatment .................................................. 2-100Appendix A, References .................................... 1-4Appendix B, Abbreviations and Terms ............ 1-4Application of External Power ...................... 2-101Approved Fuels .............................................. T2-10Approved Military Fuels, Oil, Fluids

and Unit Capacities ...................................... T2-9Army Aviation Safety Program ........................ 1-7Arrival Briefing ................................................ 8-75Audio Control Panels .............................. 3-6, F3-1Autoignition System ........................................ 2-30Automatic Flight Control System .................... 3-29Automatic Direction Finder

(DF-203) ............................................ 3-27, F3-18Autopilot Controller ...................................... F3-23

SubjectParagraph, Figure,

Table Number

A

Autopilot System Limits .................................. T3-1Autopilot Limitations .......................................... 5-9Autopilot Approaches ...................................... 8-61Avionics Equipment Configuration .................... 3-2

B

Bailout .............................................................. 9-31Balance Definitions ............................................ 6-8Bank and Pitch Limits .................................... 5-27Before Exterior Check .................................. 8-12Before Landing ................................................ 8-47Before Leaving Aircraft .................................. 8-53Before Starting Engines .................................. 8-29Before Takeoff ................................................ 8-39Before Taxiing .................................................. 8-36Brake Deice Limitations .................................. 5-11Brake Deice System ........................................ 2-56

C

Cabin and Cargo Doors .................................. F2-9Cabin Door Caution Light Illuminated ............ 9-14Cabin Pressure Limits .................................... 5-33Caution/Advisory Annunciator Panel

Legend .......................................................... T2-7Center of Gravity Limitations ................ 5-17, 6-13Center Section, Right Side ............................ 8-21Center Section Left Side ................................ 8-18Chart C - Basic Weight and

Balance Record .................................... 6-9, F6-2Charts and Forms .............................................. 6-5Checklist ............................................................ 8-9Checks (Checklist Symbols) ............................ 8-11Chemical Toilet ................................................ 2-66Chip Detector Warning Light

Illuminated ...................................................... 9-9Cigarette Lighters and Ash Trays .................. 2-65Class (Aircraft) .................................................. 6-2Climb ................................................................ 8-43Cockpit ............................................................ F2-8

Index-1

TM 55-1510-221-10


Table Number

C

Cold Weather Operations ................................ 8-69Comments Pertinent to the Use of

Performance Graphs .................................... 7-16Condition Levers .............................................. 2-23Conditions (At Stapleton

International) .................................................. 7-2Control Pedestal .............................................. F2-7Control Wheels .................................... 2-37, F2-17Copilot’s Encoding Altimeter .......................... 2-82Copilot’s Horizontal Situation

Indicator ...................................................... F3-11Copilot’s Encoding Altimeter .............. 3-34, F3-29Copilot’s Gyro Horizon Indicator ........ 3-20, F3-13Cracked Cabin Window (Inner Panel) ............ 9-27Cracked Cabin Window (Outer Panel) .......... 9-26Cracked Cabin Window/Windshield .............. 5-34Cracked Windshield ........................................ 9-26Crew Briefings .......................................... 8-6, 8-73Crossfeed Fuel Flow ...................................... F2-15Crosswind Limitation ...................................... 5-31Cruise .............................................................. 8-44Cylinder Capacity vs Pressure

and Temperature ........................................ F2-20

D

Definition of Landing Terms ............................ 9-3Defrosting System .......................................... 2-50Density Variation of Aviation Fuel .................. F6-4Departure Briefing .......................................... 8-74Descent ............................................................ 8-45Descent-Arrival ................................................ 8-46Description (Electrical System) ...................... 2-73Description (Navigation) .................................. 3-16Description (Communications) .......................... 3-4Description (Manual) .......................................... 1-3Description (Propellers) .................................. 2-42Description (Flight Controls) .......................... 2-36Description (Emergency Equipment) .............. 2-12Desert Operation and Hot Weather

Operation ...................................................... 8-70Destruction of Army Materiel to

Prevent Enemy Use ........................................ 1-8Dimensions (Aircraft) ........................................ 2-3

Subject

D

Paragraph, Figure,Table Number

Ditching .................................................. 9-29, F9-2Diving ................................................................ 8-64Draining Moisture from Fuel System .............. 2-92Duct Overtemp Caution Annunciator

Light Illuminated .......................................... 9-10DC Electrical System .................................... F2-22DC Power Supply ............................................ 2-74

E

Electrical System Emergencies ...................... 9-22Emergency Entrance .......................................... 9-6Emergency Exits and Equipment ............ 9-5, F9-1Emergency Body Positions ............................ F9-3Emergency Descent ........................................ 9-23Emergency Locator Transmitter (ELT) 3-15, F3-8Emergency Lighting ........................................ 2-78Empennage, Area 5 ........................................ 8-26Engine Bleed Air System Failure .................... 9-12Engine Chip Detection System ...................... 2-28Engine Clearing ................................................ 8-33Engine Compartment Cooling ........................ 2-17Engine Fire Extinguisher System .......... 2-26, T2-1Engine Fire Detection System ........................ 2-25Engine Fuel Control System .......................... 2-21Engine Ice Protection Systems ...................... 2-20Engine Ignition System .................................... 2-29Engine Instruments .......................................... 2-32Engine Limitations .......................................... 5-13Engine Malfunction ............................................ 9-7Engine Runup .................................................. 8-38Engine Shutdown ............................................ 8-52Engine Starter-Generators .............................. 2-31Engines ............................................................ 2-16Entrance and Exit Provisions .......................... 2-10Environmental Controls .................................. 2-72Environmental Systems ................................ F2-21Exceeding Operational Limits .......................... 5-3Exhaust and Propeller Danger Area .............. F2-5Exhaust Danger Area ........................................ 2-6Explanation of Change Symbols .................... 1-10Extent of Coverage (Weight, Balance

and Loading) .................................................. 6-1Exterior Inspection .......................................... F8-1

Index-2

TM 55-1510-221-10

Subject

E


Exterior Check. 8-13Exterior Lighting. . . . . . . . . . . . . . . . . . . . . . . . . 2-76.F2-27

F

Feathering Provisions . . . . . . . . . . . . . . . . . . . . . . . . . . 2-43Ferry Chair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-64Ferry Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-35Filling Fuel Tanks. . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-91Fire.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-20First Aid Kits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-13First Engine Start (Battery Start) . . . . . . . . . . . . . . . . . .8-30First Engine Start (GPU Start) . . . . . . . . . . . . . . . . . . . 8-34Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-66Flight Controls Lock . . . . . . . . . . . . . . . . . . . . . 2-39. F2-18Flight Controls Malfunction . . . . . . . . . . . . . . . . . . . . .9-30Flight Director. . . . . . . . . . . . . . . . . . . . . . . . . F3-21,F3-22Flight Envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F5-2Flight Plan. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-5Flight Planning. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .7-12Flight Under IMC (Instrument

Meteorological Conditions. . . . . . . . . . . . . . . . . . . . 5-30Foreign Object Damage Control . . . . . . . . . . . . . . . . . .2-19Forms and Records . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-9Friction Lock Knobs . . . . . . . . . . . . . . . . . . . . . . . . . . .2-24Fuel and Oil Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-12Fuel Handling Precautions. . . . . . . . . . . . . . . . . . . . . . .2-90F u e l L o a d . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-11Fuel Management Panel . . . . . . . . . . . . . . . . . . . . . . F2-14Fuel Quantity Data. . . . . . . . . . . . . . . . . . . . . . . . . . . . T2-2Fuel Sample . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-14Fuel Sump Drain Locations . . . . . . . . . . . . . . . . . . . . . T2-3Fuel Supply System. . . . . . . . . . . . . . . . . . . . . . . . . . . .2-33Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .9-21Fuel System Anti-Icing. . . . . . . . . . . . . . . . . . . . . . . . .2-57Fuel System Limits. . . . . . . . . . . . . . . . . . . . . . . . . . . .5-10Fuel System Management . . . . . . . . . . . . . . . . . . . . . . .2-34Fuel System Schematic . . . . . . . . . . . . . . . . . . . . . . . . F2-13Fuel Types . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-93Fuselage. Left Side, Area 6 . . . . . . . . . . . . . . . . . . . . . .8-27Fuselage, Right Side, Area 4 . . . . . . . . . . . . . . . . . . . . .8-25Fuselage. Underside. . . . . . . . . . . . . . . . . . . . . . . . . . . .8-19

Subject

G


General (Aircraft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2General (Instrument Flight) . . . . . . . . . . . . . . . . . . . . . .8-54General (Introduction) . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-1General (Operating Limits) . . . . . . . . . . . . . . . . . . . . . . .5-2General (Servicing, Parking and Mooring) . . . . . . . . . .2-89General Exterior Arrangement . . . . . . . . . . . . . . . . . . . F2-1General Interior Arrangement . . . . . . . . . . . . . . . . . . . .F2-2Generator Limits . . . . . . . . . . . . . . . . . . . . . . . . .5-16. T5-2Go-Around. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-50Gravity Feed Fuel Flow . . . . . . . . . . . . . . . . . . . . . . . F2-16Ground Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-103Ground Turning Radius . . . . . . . . . . . . . . . . . . . . .2-4,F2-4Gyromagnetic Compass Systems . . . . . . . . . . . . . . . . . 3-22

H

HF Communication Set (KHF-950). . . . . . . . . . . . . . . .3-14HFControl Panel (718 U-5). . . . . . . . . . . . . . . . . . . . .F3-7Hand-Operated Fire Extinguisher . . . . . . . . . . . . . . . . .2-14Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-69Horizontal Situation Indicators . . . . . . . . . . . . . . . . . . .3-18

I

Ice and Rain (Typical) . . . . . . . . . . . . . . . . . . . . . . . . . .8-72Icing (Severe). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-72AIcing Limitations (Typical) . . . . . . . . . . . . . . . . . . . . 5-30AIcing Limitations (Severe) . . . . . . . . . . . . . . . . . . . . . 5-30BIce Vane Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-11Immediate Action Emergency Checks . . . . . . . . . . . . . . 9-2I n d e x . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l- 1-6Inertial Navigation System . . . . . . . . . . 3-30. F3-24, F3-25Inflating Tires. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-97Installation of Protective Covers . . . . . . . . . . . . . . . . .2-105Instrument Approaches. . . . . . . . . . . . . . . . . . . . . . . . .8-60Instrument Climb. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-57Instrument Cruise. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-58Instrument Descent. . . . . . . . . . . . . . . . . . . . . . . . . . . .8-59Instrument Flight Procedures . . . . . . . . . . . . . . . . . . . . 8-55Instrument Landing System Limits . . . . . . . . . . . . . . . .5-36Instrument Marking Color Codes . . . . . . . . . . . . . 5-6. F5-1Instrument Markings, . . . . . . . . . . . . . . . . . . . . . . . . . . .5-5Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-29Instrument Takeoff. . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-56Intentional Engine Out Speed . . . . . . . . . . . . . . . . . . . ,5-37Interior Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .8-28

Change 5 Index-3

TM 55-1510-221-10


Table Number

I

Interior Lighting .............................................. 2-77Introduction (Aircraft & Systems

Description and Operation) .......................... 2-1Introduction (Avionics) ...................................... 3-1Introduction to Performance ............................ 7-1Introduction (Adverse Environmental

Conditions) .................................................... 8-68

L

Landing ............................................................ 8-49Landing on Unprepared Runway .................... 5-38Landing with Inoperative

Wing Flaps (UP) .......................................... 9-25Landing Emergencies ...................................... 9-24Landing Gear Extension Speed ...................... 5-21Landing Gear System ........................................ 2-7Landing Gear Retraction Speed .................... 5-22Landing Information ........................................ 7-15Left Engine and Propeller .............................. 8-17Left Main Landing Gear .................................. 8-16Left Wing, Area 1 ............................................ 8-15Level Flight Characteristics ............................ 8-67Line Up ............................................................ 8-40Load Planning .................................................. 6-14Loading Procedure .......................................... 6-15Loss of Pressurization

(Above 10,000 Feet) .................................... 9-13Low Oil Pressure .............................................. 9-8

M-130 Flare and Chaff DispensingSystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3, F4-2, F4-3

Malfunction Indications andProcedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T3-4

Malfunction Code Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T3-3Maneuvering Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-65Maneuvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-26Marker Beacon Receiver . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-26Marker Beacon Audio Control Panel . . . . 3-7, F3-2Maximum Allowable Airspeed . . . . . . . . . . . . . . . . . . . . . . . . 5-20Maximum Design Maneuvering Speed . . . . . . . . . . 5-25

Subject

M


Maximum Design Sink Rate ............................ 5-35Maximum Glide ................................................ F9-2Maximum Takeoff Weight Permitted

By Enroute Climb .......................................... 7-5Maximum Weights ............................................ 2-5Microphones, Switches and Jacks .................. 3-5Minimum Crew Requirements .......................... 5-4Minimum Oil Temperature Required

For Flight ............................................ 5-39, F5-3Minimum Single-Engine Control

Airspeed (V,,) .............................................. 5-24Miscellaneous Instruments .............................. 2-88Mission Avionics Coverage .............................. 4-1Mission Control Panel Annunciator

Legend .......................................................... T2-8Mission Control Panel ............................ 4-2, F4-1Mission Equipment ........................................ F2-25Mission Planning ................................................ 8-1Mooring .............................................. 2-106, F2-33

N

Nose Section, Area 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-20NAV 1 - NAV 2 Control Panel .. . . . . . . .. . . .. . . ... . .. F3-17

O

Obstacle Clearance Approach andMinimum Run Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-48

Occupants Useful Load ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T6-1Oil Supply System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-27Operating Procedures and Maneuvers . . . . . . . . . . . . 8-7Operating Limits and Restrictions . . . . . . . . . . . . . . . . . . . . 8-2Outside Air Temperature (OAT) Gage . . . . . . . . . . . . 2-86Overhead Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F2-12Overhead Circuit Breaker Panel . . . . . . . . . . . . . . . . . . F2-26Over-temperature and Overspeed

Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-14, T5-1Oxygen Cylinder Capacity Example

Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-62Oxygen Duration Example Problem . . . . . . . . . . . . . . 2-61Oxygen Duration In Minutes

Foot System ............................................... T2-5

Index-4

TM 55-1510-221-10


Table Number

O

Oxygen Flow Planning Ratesvs Altitude .................................................... T2-4

Oxygen Requirements .................................... 5-32Oxygen System ................................................ 2-60Oxygen System Schematic .......................... F2-19Oxygen System Servicing Pressure .............. F2-31

P

Parking .......................................................... 2-104Parking Brake .................................................... 2-9Parking, Covers, Ground Handing,

and Towing ................................................ F2-32Performance ...................................................... 8-4Performance Example ...................................... 7-4Pilot’s .............................................................. F3-10Pilot’s Altimeter Indicator ........ 2-83, 3-33, F3-28Pilot’s Attitude Director Indicator ...... 3-19, F3-12Pitot and Static System ...................... 2-79, F2-28Pitot and Stall Warning Heat System ............ 2-54Pitot Heat Limitations ...................................... 5-12Placard Items .................................................. 1-13Power Definitions for Engine Operations ...... 5-15Power Levers .................................................. 2-22Power Source .................................................... 3-3Pressure Altitude .............................................. 7-3Pressurization System .................................... 2-59Principal Dimensions ...................................... F2-3Propeller Test Switches .................................. 2-45Propeller Governors ........................................ 2-44Propeller Tachometers .................................... 2-49Propeller Reversing ........................................ 2-48Propeller Levers .............................................. 2-47Propeller Synchrophaser ................................ 2-46Propeller Electrothermal Anti-Ice

System .......................................................... 2-53Propeller Limitations .......................................... 5-7Propeller Failure (Over 2080 RPM) ................ 9-19Purpose (Operating Limits and

Restrictions) .................................................... 5-1Purpose (Weight and Balance) ........................ 6-4PT6A-41 Engine .................. . . ....................... F2-11


Table Number

Radar Signal Detecting Set(AN/APR-39(V)1) ........................ 4-6, F4-4, F4-5

Radar Warning Receiver(AN/APR-44() (V3) ................................ 4-7, F4-6

Radio Altimeter Indicator .................... 3-24, F3-16Radio Magnetic Indicators (RMI) .......... 3-17, F3-9Recommended Fluid Dilution Chart ............ T2-12Relief Tube ...................................................... 2-68Required Equipment Listing .................. 5-40, F5-4Reserve Fuel .................................................... 7-13Responsibility (Wgt and Balance) .................... 6-6Right Engine and Propeller ............................ 8-22Right Main Landing Gear ................................ 8-23Right Wing, Area 3 .......................................... 8-24Rudder System ................................................ 2-38

S

Seats .................................................. 2-11, F2-10Second Engine Start (GPU Start) .................. 8-35Second Engine Start (Battery Start) .............. 8-31Securing Loads ................................................ 6-16Servicing Oil System ...................................... 2-95Servicing ........................................................ F2-30Servicing Oxygen System ............................ 2-102Servicing the Air Conditioning

System .......................................................... 2-99Servicing the Chemical Toilet ........................ 2-98Servicing Hydraulic Brake System

Reservoir ...................................................... 2-96Single Phase AC Electrical System .............. F2-23Single-Engine Go-Around ................................ 9-18Single-Engine Landing Check ........................ 9-17Single-Engine Before Landing ........................ 9-16Single-Engine Descent/Arrival ........................ 9-15Spins ................................................................ 8-63Stall Warning System ...................................... 2-55Stalls ...................................................... 8-62, F8-2Standard Alternate and Emergency

Fuels ............................................................ T2-11Standby Magnetic Compass .......................... 2-87Starter Limitations ............................................ 5-8Subpanels ........................................................ F2-6Sun Visors ........................................................ 2-67

Index-5

TM 55-1510-221-10


Table Number

S

Surface Deicing System .................................. 2-51Survival Kits .................................................... 2-15System Daily Preflight/Re-Arm Test ................ 4-4

T

Takeoff .............................................................. 8-41Takeoff Climb Gradient - One

Engine Inoperative ...................................... 7-11Takeoff Distance (Flaps 0%) ............................ 7-8Takeoff Flight Path Example .................. 7-9, F7-1Takeoff Weight to Achieve Positive ................ 7-6Taxiing .............................................................. 8-37Temperature Limits .......................................... 5-29Three Phase AC Electrical System .............. F2-24Transponder Set (APX-100) ................ 3-32, F3-27Trim Tabs ........................................................ 2-40Turbulence and Thunderstorm Operation ...... 8-71Turn and Slip Indicators .......... 2-80, 3-21, F3-14TACAN Systems ...................... 3-28, F3-19, F3-20

U

Unpressurized Ventilation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-71Use of Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-10Use of Fuels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-94Use of Words Shall, Will, Should,

and May . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-12UHF Command Set (AN/ARC-164) . . . . . . . . . . 3-8, F3-3


Table Number

V

Various Values for UTM GridCoefficients . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . T3-2

Vertical Velocity Indicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-84Voice Order Wire (AN/ARC-164) . . . . . . . . . . . . . . . . . . . . . . 3-9Voice Security System TSEC/KY-58

(Provisions Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-13Voice Security System TSEC/KY-28

(Provisions Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-12, F3-6VHF AM-FM Command Set

(AN/ARC- 186) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-11, F3-5VHF-AM Communications (VHF-2OB) . . 3-10, F3-4VOR/LOC Navigation System . . . . . . . . . . . . . . . . . . . . . . . . 3-25

W

Warning Annunciator Panel Legend . . . . . . . . . . . . . . T2-6Warnings, Cautions, and Notes . . . . . . . . . . . . . . . . . . . . . . . . 1-2Weather Radar Set (AN/APN-215) . . . . 3-31, F3-26Weight and Balance Clearance

Form F, DD Form 365-4 . . . . . . . . . . . . . . . . . . . . . . 6-9, F6-3Weight Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-7Weight Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-18Weight, Balance, and Loading . . . . . . . . . . . . . . . . . . . . . . . . . . 8-3Wind Swell Ditch Heading

Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . F9-4Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10Windshield Wipers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-63Windshield Electrothermal

Anti-Ice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-58Wing Flap Extension Speeds . . . . . . . . . . . . . . . . . . . . . . . . . . 5-23Wing Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-41

Zero Fuel Weight Limitation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-14

Index-6

TM 55-1510-221-10


Official:

CARL E. VUONOGeneral, United States Army

Chief of Staff

WILLIAM J. MEEHAN IIBrigadier General, United States Army

The Adjutant General

DISTRIBUTION:To be distributed in accordance with DA Form 12-31, -10 & CL Maintenance require-

ments for RC-12H Airplane, Reconnaissance.

* U.S. GOVERNMENT PRINTING OFFICE : 1994 0 - 300-769 (12205)

*U.S. GOVERNMENT PRINTING OFFICE 1994-300-769-12205

These are the instructions for sending an electronic 2028The following format must be used if submitting an electronic 2028. The subject line must be exactly thesame and all fields must be included; however only the following fields are mandatory: 1, 3, 4, 5, 6, 7, 8,9, 10, 13, 15, 16, 17, and 27.

From: ‘Whomever” <[email protected] .mil>To: [email protected] .mil

Subject: DA Form 20281. From: Joe Smith2. Unit: home3. Address: 4300 Park4. City: Hometown5. St: MO6. 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: 926. Total: 12327. Text:This is the text for the problem below line 27.

PIN: 065798-000

This fine document...

Was brought to you by me:

Liberated Manuals -- free army and government manuals

Why do I do it? I am tired of sleazy CD-ROM sellers, who take publicly available information, slap “watermarks” and other junk on it, and sell it. Those masters of search engine manipulation make sure that their sites that sell free information, come up first in search engines. They did not create it... They did not even scan it... Why should they get your money? Why are not letting you give those free manuals to your friends?

I am setting this document FREE. This document was made by the US Government and is NOT protected by Copyright. Feel free to share, republish, sell and so on.

I am not asking you for donations, fees or handouts. If you can, please provide a link to liberatedmanuals.com, so that free manuals come up first in search engines:

<A HREF=http://www.liberatedmanuals.com/>Free Military and Government Manuals</A>

– SincerelyIgor Chudovhttp://igor.chudov.com/

– Chicago Machinery Movers

http://www.liberatedmanuals.com/

https://www.machinerymoverschicago.com/

http://igor.chudov.com/

http://www.liberatedmanuals.com/

Date post:	06-Mar-2021
Category:	Documents
Upload:	others
View:	4 times
Download:	0 times

HEADQUARTERS, DEPARTMENT · URGENT CHANGE NO. 5 TM 55-1510-221-10 C5 HEADQUARTERS DEPARTMENT OF THE...

Documents