Bell Uh-1h-II Helicopter Flight Manual

BHT PUB-92-004-10

TECHNICAL MANUAL

OPERATOR’S MANUAL BELL MODEL UH-1H-II HELICOPTER

BHT PUB-92-004-10

TECHNICAL MANUAL

OPERATOR’S MANUAL BELL MODEL UH-1H-II HELICOPTER

COPYRIGHT NOTICE

2002 COPYRIGHT BELL ® HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON, A DIVISION OF TEXTRON CANADA LTD. ALL RIGHTS RESERVED

18 MARCH 1994 REVISION 7 — 1 JULY 2002

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PROPRIETARY RIGHTS NOTICE These data are proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than as a guide for helicopter repair is forbidden without prior written authorization from Bell Helicopter Textron.

Additional copies of this publication may be obtained by contacting: Commercial Publication Distribution Center Bell Helicopter Textron Inc. P. O. Box 482 Fort Worth, Texas 76101

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LIST OF EFFECTIVE PAGES Insert latest revision pages as indicated: dispose of superseded pages. On a revised page the portion of the text and illustration affected by the latest technical revision is indicated by a black bertical line. Revised pages without a black vertical line are a result of text shift from page to page or non-technical corrections. Original ................... 0................ 18 Mar 94 Revision................... 1................15 May 97 Revision................... 2................. 12 Jan 98 Page No.

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Cover ................................................... 7 Title ..................................................... 7 P/N....................................................... 0 A/B....................................................... 7 Warning 1 — Warning 2 .................... 0 i............................................................ 5 ii........................................................... 7 1-1/1-2 Blank ...................................... 1 2-1 — 2-44 ........................................... 7 2-45/2-46 Blank .................................. 7 3-1........................................................ 1 3-2 — 3-8 ............................................. 0 3-9........................................................ 1 3-10...................................................... 0 3-11...................................................... 1 3-12...................................................... 0 3-13...................................................... 1 3-14...................................................... 0 3-15 — 3-17 ......................................... 0 3-18...................................................... 1 3-19 — 3-22 ......................................... 0 4-1........................................................ 6 4-2........................................................ 0 4-2A — 4-2B........................................ 6 4-3........................................................ 6 4-4........................................................ 5 4-5/4-6 Blank ...................................... 5 5-1........................................................ 5 5-2........................................................ 6 5-3 — 5-4 ............................................. 7 5-5........................................................ 6 5-6........................................................ 0 5-6A/5-6B Blank ................................. 6 5-7........................................................ 0 5-8........................................................ 7 5-9........................................................ 1 5-10...................................................... 0 5-11/5-12 Blank .................................. 0 6-1 — 6-6 ............................................. 0 6-7........................................................ 1 6-8 — 6-14 ........................................... 0 6-15...................................................... 6 6-16...................................................... 0 6-17...................................................... 1 6-18...................................................... 0 6-19...................................................... 6 6-20...................................................... 2 6-21...................................................... 6 6-22 — 6-24 ......................................... 0 6-25...................................................... 5

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8-16 ..................................................... 6 8-17 ..................................................... 6 8-18 ..................................................... 1 8-19 ..................................................... 5 8-20 — 8-21......................................... 0 8-22 — 8-25......................................... 5 8-26 ..................................................... 7 8-27/8-28 Blank .................................. 5 9-1 ....................................................... 5 9-2 — 9-4............................................. 0 9-5 ....................................................... 7 9-6 ....................................................... 5 9-7 — 9-8............................................. 1 9-9 — 9-10........................................... 5 9-11 ..................................................... 7 9-12 ..................................................... 5 9-13/9-14 Blank .................................. 7 Index 1 — Index 6 .............................. 0 Index 7/Index 8 Blank........................ 3

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Personnel performing operations, procedures and practices which are included or implied in this technical manual shall observe the following warnings. Disregard of these warnings and precautionary information can cause serious injury or DEATH, or an aborted mission.

STARTING ENGINES Coordinate all cockpit actions with ground observer. Insure that rotors and blast areas are clear and fire guard is posted.

FIRE EXTINGUISHER Exposure to high concentrations of monobromotrifluoromethane (CF3Br) extinguishing agent or decomposition products should be avoided. The liquid should not be allowed to come into contact with the skin, as it may cause frostbite or low temperature burns.

GROUND OPERATION Engines will be starter and operated only by authorized personnel. Reference AR 95–1.

ACIDS Battery electrolyte is harmful to the skin and clothing. Neutralize any spilled electrolyte by flushing contact areas thoroughly with water.

CARBON MONOXIDE When smoke, suspected carbon monoxide fumes, or symptoms of anoxia exist, the crew should immediately ventilate cabin and shut off heater.

HANDLING FUEL AND OILS Turbine fuels and lubricating oil contain additives which are poisonous and readily absorbed through the skin. Do not allow them to remain on skin longer than necessary.

Warning 1

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RADIO ACTIVE MATERIALS Self–luminous dials and ignition units contain radioactive materials. If such an instrument or unit is broken or becomes unsealed, avoid personal contact.

WARNING Sound pressure levels in the aircraft during some operating conditions exceed the Surgeon General’s hearing conservation criteria as defined in TB MED 251. Hearing protection devices, such as the aviator helmet or ear plugs are required to be worn by all personnel in and around the aircraft during its operation.

Warning 2

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TABLE OF CONTENTS Chapter/Section

Page

CHAPTER 1

INTRODUCTION ..............................................................................................1-1

CHAPTER 2 Section I II III IV V VI VII VIII IX X XI XII XIII XIV XV

HELICOPTER AND SYSTEMS DESCRIPTION AND OPERATION Helicopter .......................................................................................................... 2-1 Emergency equipment .................................................................................... 2-10 Engine and related systems ........................................................................... 2-10 Helicopter fuel system .................................................................................... 2-22 Flight control system ...................................................................................... 2-31 Hydraulic system ............................................................................................ 2-32 Power train systems ....................................................................................... 2-33 Main and tail rotor group ............................................................................... 2-34 Utility system.................................................................................................. 2-34 Heating and ventilation ................................................................................. 2-35 Electrical power supply and distribution system.......................................... 2-35 LIghting........................................................................................................... 2-37 Flight instruments.......................................................................................... 2-39 Servicing, parking and mooring .................................................................... 2-42 Auxiliary equipment ....................................................................................... 2-43

CHAPTER 3 Section I II III

AVIONICS Communications ............................................................................................... 3-1 Navigation ....................................................................................................... 3-11 Transponder and Radar.................................................................................. 3-20

CHAPTER 4

MISSION EQUIPMENT .................................................................................. 4-1

CHAPTER 5 Section I II III IV V VI VII VIII

OPERATING LIMITS AND RESTRICTIONS General ..............................................................................................................5-1 System limits .................................................................................................... 5-1 Power limits ...................................................................................................... 5-4 Loading limits ................................................................................................... 5-5 Airspeed limits .................................................................................................. 5-8 Maneuvering limits........................................................................................... 5-9 Environmental restrictions .............................................................................. 5-9 Height velocity .................................................................................................. 5-9

CHAPTER 6 Section I II III IV V VI VII

WEIGHT/BALANCE AND LOADING General ...............................................................................................................6-1 Weight and balance ...........................................................................................6-1 Personnel............................................................................................................6-9 Mission equipment (Not applicable) Cargo loading .................................................................................................. 6-15 Fuel/oil............................................................................................................. 6-21 Allowable loading............................................................................................ 6-21

CHAPTER 7 Section I II

PERFORMANCE DATA Introduction....................................................................................................... 7-1 Torque available ............................................................................................... 7-7

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TABLE OF CONTENTS Chapter/Section

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Page III IV V VI VII VIII IX X XI XII XIII XIV XV

Hover ................................................................................................................ 7-9 Control Margin................................................................................................ 7-15 Takeoff ............................................................................................................. 7-19 Cruise .............................................................................................................. 7-25 Drag ................................................................................................................. 7-77 Climb performance/climb — descent ............................................................. 7-83 Fuel flow .......................................................................................................... 7-87 Nautical miles and time and range ............................................................... 7-90 Autorotational glide characteristics .............................................................. 7-94 Airspeed limits ................................................................................................ 7-96 Airspeed correction ......................................................................................... 7-98 Temperature.................................................................................................. 7-100 Performance planning .................................................................................. 7-102

CHAPTER 8 Section I II III IV V VI

NORMAL PROCEDURES Mission planning .............................................................................................. 8-1 Operating procedures and maneuvers............................................................. 8-1 Instrument flight ............................................................................................ 8-15 Flight characteristics ..................................................................................... 8-16 Adverse environmental conditions................................................................. 8-22 Crew duties .................................................................................................... 8-27

CHAPTER 9 Section I

EMERGENCY PROCEDURES Helicopter systems............................................................................................ 9-1

INDEX

Index ........................................................................................................... Index-1

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CHAPTER 1 INTRODUCTION

IMPORTANT In order to obtain complete information and derive maximum benefits from this manual, it is necessary to read this chapter carefully and thoroughly.

1-1.

PURPOSE.

The purpose of this manual is to supply you with the latest information and performance data derived from flight test programs and operational experiences. The study and use of this manual will enable you to perform the assigned missions and duties with maximum efficiency and safety. Your ability and experience are recognized. It is not the function of this manual to teach the pilot how to fly; basic flight principles and elementary instructions are not included. The contents of this manual will provide you with a general knowledge of Bell UH1H-II including general flight characteristics and specific normal and emergency operating procedures.

An operating procedure, practice, etc., which, if not correctly followed, could result in personnel injury or loss of life.

An operating procedure, practice, etc., which, if not strictly observed, could result in damage to or destruction of equipment.

a. Reports. Reports necessary to comply with the Army Safety Program are prescribed in detail in AR 385-40.

NOTE

b. Forms. DA forms and procedures used for equipment maintenance will be only those prescribed by TM 38-750 and 55-405-9.

1-2. WARNINGS, CAUTIONS, AND NOTES, DEFINITION. Warnings, Cautions, and Notes are used to emphasize important and critical instructions and are used for the following conditions.

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

1-3.

REPORTING OF IMPROVEMENTS.

Report of error, omissions, and recommendations for improving this publication by the individual user is encouraged. Reports shall be submitted to the cognizant authority.

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CHAPTER 2 HELICOPTER AND SYSTEMS DESCRIPTION AND OPERATION Section I. HELICOPTER 2-1.

GENERAL DESCRIPTION.

2-3. PRINCIPAL DIMENSIONS.

The Bell UH-1H-II helicopter is a thirteen-place single engine helicopter. The maximum gross weight for takeoff is 10,500 pounds. (Figure 2-0).

Figure 2-2 depicts the principal dimensions.

2-4. TURNING RADIUS. Figure 2-3 depicts the minimum turning radius.

2-2. GENERAL ARRANGEMENT. Figure 2-1 depicts the general arrangement. Indexed items include access openings and most of the items referred to in the exterior check paragraph in Section II of Chapter 8.

2-5. FUSELAGE. The fuselage is the forward section of the airframe extending from the nose to the forward end of

Figure 2-0. Bell model UH-1H-II

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the tailboom. The fuselage consists primarily of two longitudinal beams with transverse bulkheads and metal covering. The main beams are the supporting structure for the cabin, landing gear, fuel tanks, transmission, engine, and tailboom. The external cargo suspension unit is attached to the main beams near the CG of the helicopter.

b. Cargo/Passenger Doors. The two cargo/ passenger doors are formed aluminum frames with transparent plastic windows in the upper section (figure 2-1). These doors are on rollers and slide aft to the open position allowing full access to the cargo doors. They provide a larger entrance to the cargo area.

2-6. TAILBOOM.

2-10. PILOT/COPILOT SEAT.

The tailboom section is bolted to the aft end of the fuselage and extends to the aft end of the helicopter. It is a tapered, semi-monocoque structure employing magnesium skins, aluminum longerons, and stringers. The tailboom supports the tail rotor, vertical fin, and synchronized elevator. It houses the tail rotor driveshaft, some electronic equipment, and a baggage compartment (figure 2-1). Baggage compartment is located in forward end of tailboom and has a capacity of 28 cubic feet (0.8 cubic meter). Compartment can carry up to 400 pounds (181 kilograms) of baggage or other cargo, which can be secured using tiedown fittings provided. Access door is on right side of tailboom and is provided with a key lock for security of compartment contents. Two interior lights illuminate when door is open. DOOR LOCK caution light, on caution panel, illuminates when door is open or is not properly latched. A smoke detector is installed in compartment and is connected to BAGGAGE FIRE warning light located on instrument panel.

2-7. LANDING GEAR SYSTEM. a. Main Landing Gear. The main landing gear consists of two aluminum arched crosstubes mounted laterally on the fuselage with two longitudinal skin tubes attached to the crosstubes. The skid tubes are made of aluminum and have steel skid shoes attached to the bottom to minimize skid wear. b. Tail Skid. A tubular steel tail skid is installed on the aft end of the tailboom. It acts as a warning to the pilot upon an inadvertent tail-low landing.

2-8. Crew Compartment Diagram. The pilot and copilot compartment is depicted in figure 2-4.

2-9. Crew and Cargo Doors. a. Pilot and Copilot Doors. The pilot and copilot doors (figure 2-1) are formed aluminum frames with transparent plastic windows in the upper section. Ventilation is supplied by the sliding panels in the windows. Cam-type door latches are used and doors are equipped with jettisonable door releases. 2-2

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a. Pilot and Copilot Seat. Armored seats are installed in the helicopter for the pilot and copilot. They are equipped with lap safety belt and inertia– reel shoulder harness. They are adjustable fore and aft and vertically. The vertical adjustment handle is under the right side of the seat and the fore and aft handle is on the left. The seats are equipped with a quick release, on each side at the back of the seat, for reclining the seat to aid in removal of injured pilot/copilot. The seat back, bottom, and sides are protected by ceramic and aluminum armor plate. Hip and shoulder areas are protected by ceramic type armor.

Upward force of springs located on seats may cause rapid vertical movement. b. Inertial Reel Shoulder Harness. An inertial reel and shoulder harness is incorporated in the pilot and copilot seats with manual lock–unlock handle (figures 2-4 and 2-5). The control handles are located on the side of the seat. With the control in the unlocked position and the shoulder straps properly adjusted, the reel strap will extend to allow the occupant to lean forward; however, the reel automatically locks when the helicopter encounters an impact force of 2 to 3 “G” deceleration. The reel can be locked from any position and will take up slack in harness. To release the lock it is necessary to lean back slightly to release tension on the lock and move the control handle to the unlock position. It is possible to have pressure against the seat back whereby no additional movement is possible and the lock cannot be released. If this condition occurs, it will be necessary to loosen the harness. The reel should be manually locked for emergency landing. NOTE Straps must be adjusted to fully retract within the inertia reels to prevent dynamic overshoot in the event of impact. NOTE Seat belt must be securely fastened and firmly tightened prior to adjustment of shoulder harness to prevent submarining in event of impact.

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Figure 2-1. General arrangement (Sheet 1 of 2)

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Figure 2-1. General arrangement (Sheet 2 of 2)

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Figure 2-2. Helicopter principle dimensions

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Figure 2-3. Turning radius

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Figure 2-4. Compartment diagram (Sheet 1 of 2)

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Figure 2-4. Compartment diagram (Sheet 2 of 2)

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Figure 2-5. Pilot/Copilot seat

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installed on the instrument panel is depicted in figure 2-6.

2-11. PERSONNEL SEATS. Various arrangements of personnel seats can be installed to accommodate from one to eleven personnel besides the pilot and copilot. The seats are constructed of tubular steel and reinforced canvas. Each seat is equipped with a lap safety belt. For additional information on the personnel seats, refer to Chapter 6.

2-12. INSTRUMENTS AND CONTROLS. a. Instrument Panel. The location of all the controls, indicators, instruments, and data placards

b. Pedestal Panel. The panels and controls installed in the pedestal are depicted in figure 2-7. c. Overhead Console. The location of the controls circuit breakers installed in the overhead console is depicted in figure 2-7 and 2-11. d. Other Instruments and Controls. Instruments, controls, or indicators not shown in figures 2-6 or 27 are shown in the crew compartment diagram (figure 2-4) or in the Chapter/Section which describes their related systems.

Section II. EMERGENCY EQUIPMENT NOTE The emergency equipment location, illustration and emergency procedures are covered in Chapter 9.

2-13. PORTABLE FIRE EXTINGUISHER. A portable hand-operated fire extinguisher is carried in a bracket located on the right center door post. RefertoChapter9forillustration.

2-14. FIRST AID KITS. Four aeronautical type first aid kits have been provided in the cabin area. Two kits are secured to the right center door post. the other two kits are secured to the left center door post. First aid kits can be easily removed for immediate use. Refer to Chapter 9 for illustration of the location of the first aid kits.

Section III. ENGINE AND RELATED SYSTEMS 2-15. ENGINE. The helicopter is equipped with a T53-L-703 series engine (figure 2-8) which at 6600 rpm has the following ratings: Take-Off power TOP) — 1800 shp and Maximum Continuous Power (MCP) — 1500 shp. The helicopter is torque limited by the transmission to TOP (5 minutes) — 1290 shp and MCP — 1134 shp which corresponds to 100% and 88% on the torque gauge when operating at 6600 engine/324 rotor rpm.

2-16. ENGINE COMPARTMENT COOLING. The engine compartment is cooled by natural convection through engine compartment screens.

2-17. AIR INDUCTION SYSTEM. a. Lower and Filter Systems. The engine air inlet section draws in air through a bellmouth which is fitted with a wire screen. The bellmouth extends through the forward firewall into the air induction area. The induction area is covered by a three-piece set of air inlet filters. The filters have a doubler layer of porous foam plastic material to protect the engine from foreign matter. b. Particle Separator. This is an inertial-type separator. Particle-laden air is directed through a

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large annular chamber and through an air cleaner. A constant supply of bleed air from the engine flows through the venturi-type ejector and carries particle overboard through airframe plumbing. c. Foreign Object Damage Screen. A Foreign Object Damage (FOD) screen is installed to prevent large particles from entering the engine. NOTE The ice detector system is not applicable on helicopter equipped with the selfpurging particle separator. d. DE-ICE. Engine de-ice is a bleed air system activated by the DE-ICE switch on the ENGINE control panel (figure 2-9). In the ON position bleed air is directed through the engine inlet to provide the protection. Power losses caused when the system is on are shown in Chapter 7. In the event of dc electrical failure or when the ANTI-ICE ENG circuit breaker is out de-ice is automatically on. System power is provided by the dc essential bus and protected by the ANTI-ICE-ENG circuit breaker.

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Figure 2-6. Instrumental panel — UH-1H-II

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Figure 2-7. Pilot station diagram typical (Sheet 1 of 3)

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Figure 2-7. Pilot station diagram typical (Sheet 2 of 3)

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Figure 2-7. Pilot station diagram typical (Sheet 3 of 3)

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Figure 2-8. Engine (Sheet 1 of 2)

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Figure 2-8. Engine (Sheet 2 of 2)

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2-18. ENGINE FUEL CONTROL SYSTEM. a. Engine Mounted Components. The fuel control assembly is mounted on the engine. It consists of a metering section, a computer section and an overspeed governor. (1) The metering section is driven at a speed proportional to N1 speed. It pumps fuel to the engine through the main metering valve or, if the main system fails, through the emergency metering valve which is positioned directly by the twist grip throttle. (2) The computer section determines the rate of main fuel delivery by biasing main metering valve opening for N1 speed, inlet air temperature and pressure, and throttle position. It also controls the operation of the compressor air-bleed system, and controls operation of the variable inlet guide vanes. (3) The overspeed governor is driven at a speed proportional to N2 speed. It biases the main metering valve opening to maintain a constant selected power output shaft rpm and, in event of a power turbine overspeed, will decrease the valve opening to reduce fuel to the engine. b. Starting Fuel Flow. During engine start, energizing the starter switch opens the fuel solenoid valve, allowing fuel from the fuel regulator to flow through the starting fuel manifold and into the combustion chamber. When N1 reaches sufficient speed, the start fuel switch is de-energized, causing the solenoid valve to close and stop-starting fuel flow. Starting fuel nozzles are purged by air from the combustion chamber through a check filter valve. c. Fuel Control Switches. Fuel flow and mode of operation is controlled by switches on the pedestalmounted engine control panel (figure 2-9). The panel contains the FUEL ON/OFF switch, START FUEL switch, two INT FUEL TRANS PUMP switches, and GOV AUTO/EMER switch. The switch over to emergency fuel is accomplished by retarding the collective twist grip to idle and positioning the GOV AUTO/EMER switch to the EMER position. In the EMER position fuel is manually metered to the engine, with no automatic control features, by rotating the collective twist grip throttle. NOTE The igniter solenoid valve (engine starting fuel solenoid valve) is controlled by the engine starter switch and the igniter solenoid valve (engine starting fuel solenoid valve) cannot be individually controlled during engine starts.

d. Power Controls (Throttles). Rotating the pilot or copilot twist grip-type throttle (figure 2-4) to the full open position allows the overspeed governor to maintain a constant rpm. Rotating the throttle toward the closed position will cause the rpm to be manually selected instead of automatically selected by the overspeed governor. Rotating the throttle to the fully closed position shuts off the fuel. An idle stop is incorporated in the throttle to prevent inadvertent throttle closure. To bypass the idle detent, depress the ENGINE IDLE REL switch located on the pilots collective stick switch box and close throttle. The ENGINE IDLE REL switch is a momentary on, solenoid-operated switch. The switch receives power from the 28 Vdc bus and is protected by a circuit breaker marked IDLE STOP REL. Friction can be induced in both throttles by rotating the pilot throttle friction ring (figure 2-4) counterclockwise. The ring is located on the upper end of the pilot throttle. e. Governor Switch. The GOV switch is located on the ENGINE control panel (figure 2-9). AUTO position permits the overspeed governor to automatically control the engine rpm with the throttle in the full open position. The EMER position permits the pilot or copilot to manually control the rpm because automatic acceleration, deacceleration, and overspeed control are not provided with the GOV switch in the EMER position, control movements must be smooth to prevent compressor stall, overspeed, overtorque, or engine failure. The governor circuit receives power from the 28 Vdc essential bus and is protected by the GOV CONT circuit breaker.

2-19. ENGINE OIL SUPPLY SYSTEM. a. Description. The dry sump pressure type oil system is entirely automatic in its operation. The system consists of an engine oil tank with deaeration provisions, thermostatically controlled oil cooler with bypass valve, pressure transmitter and pressure indicator, low pressure warning switch and indicator, sight gages, oil supply return vent, and breather lines. Crashworthy breakaway valves are incorporated in the engine-to-oil tank lines. Drain valves have been provided for draining the oil tank and cooler. Pressure for engine lubrication and scavenging of return oil are provided by the enginemounted and engine-driven oil pump. The tank capacity, oil specification and grade are specified in the servicing diagram (figure 2-14).

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Figure 2-9. Engine and hydraulic control panels

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b. Oil Cooler. Engine oil cooling is accomplished by an oil cooler with thermostatic valves and bypass provisions. The cooler is housed within the fuselage area under the engine deck. Air circulation for oil cooling is supplied by a turbine fan which operates from turbine bleed air. The fan is powered at all times when the engine is operating and no control is required except the bleed air limiting orifice.

2-20. IGNITION — STARTER SYSTEM. The starter ignition switch is mounted on the pilot collective pitch control lever switch box. The switch (figure 2-4) is a trigger switch, spring-loaded to the off position. The starter and ignition unit circuits are both connected to the trigger switches. The circuits receive power from the 28 Vdc essential bus and are protected by circuit breakers marked STARTER RELAY and IGNITION SYSTEM IGNITER SOL. The starter circuit is energized when the STARTER/ GEN switch (figure 2-7) is in the START position and the trigger switch is pulled. The ignition circuit is energized when the FUEL MAIN ON/OFF switch on the engine control panel is in the ON position and the trigger switch is pulled. The ignition keylock is located by the AC circuit breaker panel.

The off position deactivates the igniters and start fuel to prevent engine starting. The on position allows engine starting.

2-21. GOVERNOR RPM SWITCH. The pilot and copilot GOV RPM INCR/DECR switches are mounted on a switch box attached to the end of the collective pitch control lever (figure 24). The switches are a three-position momentary type and are held in INCR (up) position to increase the power turbine (N2) speed or down to DECR position to decrease the power turbine (N2) speed. Electrical power for the circuit is supplied from the 28 Vdc essential bus and is protected by a circuit breaker marked GOV CONT.

2-22. DROOP COMPENSATOR. A droop compensator maintains engine rpm (N2) as power demand is increased by the pilot. The compensator is a direct mechanical linkage between the collective stick and the speed selector lever or the N2 governor. No crew controls are provided or required. The compensator will hold N2 rpm to ± 40 rpm when properly rigged. Droop is defined as the speed change in engine rpm (N2) as power is increased from a no-load condition. It is an inherent

Figure 2-10. Heating and defrosting systems controls

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Figure 2-11. DC circuit breaker panel

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Figure 2-12. AC circuit breaker panel characteristic designed into the governor system. Without this characteristic, instability would develop as engine output is increased resulting in N1 speed overshooting or hunting the value necessary to satisfy the new power condition. Design droop of the engine governor system is as much as 300 to 400 rpm (flat pitch to full power). If N2 power were allowed to droop, other than momentarily, the reduction in rotor speed could become critical. All engine instruments and indicators are mounted in the instrument panel (figure 2-6) and the pedestal (figure 2-7).

2-23. ENGINE INDICATORS.

INSTRUMENTS

AND

a. Torquemeter Indicator. The torquemeter indicator (figure 2-6) is located in the center area of the instrument panel. The indicator is connected to a transmitter which is part of the engine oil system. The torquemeter indicates torque in % torque readings of torque imposed upon the engine output shaft. The torquemeter receives power from the 28 Vdc bus and is protected by a circuit breaker marked TORQUE in the DC circuit breaker panel. b. Engine Gas Temperature (EGT) Indicator. The engine gas temperature indicator (figure 2-6) is located in the center area of the instrument panel. The indicator receives turbine gas temperature

(TGT) indications from the thermocouple probes mounted in the engine turbine inlet area. The temperature indications are in degrees celsius. The system is electrically self-generating. c. Dual Tachometer. The dual tachometer (figure 2-6) is located in the center area of the instrument panel and indicates both the engine and main rotor rpm. The tachometer inner scale is marked ROTOR and the outer scale is marked ENGINE. Synchronization of the ENGINE and ROTOR needles indicates normal operation of helicopter. The indicator receives power from the tachometer generators mounted on the engine and transmission. Connection to the helicopter electrical system is not required. d. Gas Producer Tachometer Indicator. The gas producer indicator (figure 2-6) is located in the right center area of the instrument panel. The indicator displays the rpm of the gas producer turbine speed in percent. This system receives power from a tachometer generator which is geared to the engine compressor. A connection to the helicopter electrical system is not required. e. Engine Oil Temperature and Pressure Indicator. The engine oil temperature and pressure indicator is a dual type indicator registering temperature (degrees celsius) and pressure (psig). The temperature portion receives temperature

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indications from an electrical resistance-type bulb and the pressure portion receives pressure indications from an engine oil pressure transmitter. Both temperature indicator and pressure indicator are 28 Vdc powered and circuit protection is provided by the ENG OIL PRESS circuit breaker in the DC circuit breaker panel. f. Oil Pressure Caution Light. The ENGINE OIL PRESS caution light is located in the pedestal mounted CAUTION panel (figure 2-13). The light is connected to a low pressure switch. When pressure drops below a safe limit, the switch closes an electrical circuit, causing the caution light to illuminate. The circuit receives power from the 28 Vdc essential bus and is protected by the CAUTION LIGHTS circuit breaker in the DC circuit breaker panel. g. Engine Chip Detector Caution Light. When sufficient metal particles accumulate on the magnetic plug to complete the circuit, the ENGINE CHIP DET segment illuminates (figure 2-13). The circuit receives power from the 28 Vdc essential bus and is protected by the circuit breaker marked CAUTION LIGHTS. h. Engine Oil Filter Warning Light. The ENG OIL FILTER warning light is mounted on the instrument panel (figure 2-6). When an impeding engine oil filter blockage condition exists, the ENG OIL FILTER warning light illuminates. Power is supplied from the 28 Vdc bus and circuit protection is provided by the baggage compartment light circuit breaker. i. Engine Inlet Air Filter Clogged Warning Light. The ENGINE INLET AIR warning light is mounted on the caution panel (figure 2-13). When

the inlet air filter becomes clogged, a differential pressure switch senses the condition and closes contacts to energize the filter caution light. Power is supplied from the 28 Vdc bus and circuit protection is provided by the caution light circuit breaker. j. Engine Fuel Pump Caution Light. The ENG FUEL PUMP caution light is located in the pedestal-mounted caution panel (figure 2-13). The light is connected electrically to a switch at the engine driven dual element fuel pump. Failure of either fuel pump element will close an electrical circuit illuminating the caution light. The system receives power from the 28 Vdc essential bus and is protected by a circuit breaker marked CAUTION LIGHTS. k. Emergency Fuel Control Caution Light. The emergency fuel control caution light is located in the pedestal-mounted caution panel (figure 2-13). The illumination of the worded segment GOV EMER is a reminder to the pilot that the GOV switch is in the EMER position. Electrical power for the circuit is supplied from the 28 Vdc bus and is protected by a circuit breaker marked CAUTION LIGHTS. l. Fuel Filter Caution Light. The FUEL FILTER caution light is located in the pedestal-mounted caution panel (figure 2-13). A differential pressure switch is mounted in the fuel line across the filter. When the filter becomes clogged, the pressure switch senses this and closes contacts to energize the caution light circuit. If clogging continues, the fuel bypass opens to allow fuel to flow around the filter. The circuit receives power from the 28 Vdc essential bus and is protected by a circuit breaker marked CAUTION LIGHTS.

Section IV. HELICOPTER FUEL SYSTEM 2-24. FUEL SUPPLY SYSTEM. The helicopter is equipped with a crashworthy fuel system. The fuel system consists of five interconnected cells all filled from a single fitting on the right side of the helicopter (figure 2-14). Electrically driven pumps are installed in both forward cells and provide fuel pressure to prime the fuel line to the engine driven fuel pump. Each forward fuel cell is divided into two compartments by a lateral baffle fitted with a flapper valve to allow fuel flow from front to rear. The submerged boost pump is mounted on a sump assembly near the aft end of each

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forward cell and is connected by a hose to the pressure line outlet to the engine. Part of the pump output is diverted forward through a flow switch and hose to an ejector pump at front of cell. Induced flow of the ejector pump sends fuel through a hose over the baffle into the rear part of the cell, so that no significant quantity of fuel will be unusable in any flight attitude. a. Crashworthy Fuel System. The crashworthy fuel system is designed to contain fuel during a severe, but survivable, crash impact to reduce the possibility of fire. Frangible fittings used to secure

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Figure 2-13. Caution panel

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Figure 2-14. Servicing diagram (Sheet 1 of 6)

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Figure 2-14. Servicing diagram (Sheet 2 of 6)

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Figure 2-14. Servicing diagram (Sheet 3 of 6)

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Figure 2-14. Servicing diagram (Sheet 4 of 6)

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Figure 2-14. Servicing diagram (Sheet 5 of 6)

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Figure 2-14. Servicing diagram (Sheet 6 of 6)

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the fuel cells in the airframe are designed to fail and permit relative movement of the cells, without rupture, in event of a crash; self-sealing break-away valves are installed in the fuel lines at the fuel cell outlets and certain other locations. The break-away valves are designed to permit complete separation of components without loss of fuel. Rollover vent valves are installed on the aft fuel cells to provide protection in the event of a helicoter rollover during a crash. The system has a 50 caliber ballistic protection level in the lower two-thirds of the cell. b. Closed Circuit Refueling System. The helicopter provides a closed circuit refueling system when used with the mating nozzle. This system is capable of automatic shutoff of fuel flow when full. c. Gravity Refueling NOTE When refueling with a gravity fuel nozzle a member of the flight crew should monitor the refueling operation to ensure that the receiver unit is not damaged.

If helicopter is equipped with closed circuit refueling system and fuel servicing vehicle is not equipped with related nozzle for closed circuit refueling, a gravity system may be used providing the servicing nozzle does not exceed 1.75 inches outside diameter. To refuel utilizing the gravity nozzle, it is necessary to position the inner sleeve of receiver until slot is lined up with fuel port in bottom of receiver. Position nozzle into port in order to bypass closed circuit valve. Damage could result to the closed circuit refueling system if caution is not used when fueling with the gravity servicing nozzle.

2-25. Controls and Indicators. a. Fuel switches. The fuel system switches consist of a main fuel switch, START FUEL switch, and fuel transfer switches (figure 2-9). Refer to paragraph 226a for an operational description of the fuel transfer switches. The FUEL ON/OFF switch is located on the pedestal-mounted ENGINE panel (figure 2-9). The switch is protected from being accidentally turned OFF by a spring-loaded toggle head that must be pulled up before switch movement can be accomplished. When the switch is in the ON

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position, the fuel valve opens, the electrical boost pump(s) are energized and fuel flows to the engine. When the switch is in the OFF position the fuel valve closes and the electrical boost pump(s) are deenergized. Electrical power for circuit operation is supplied by the 28 Vdc essential bus and is protected by circuit breakers FUEL VALVE LH BOOST PUMP, AND RH FUEL BOOST PUMP. b. Fuel Quantity Indicator. The fuel quantity indicator (figure 2-6) is located in the upper center area of the instrument panel. This instrument is a transistorized electrical receiver which continuously indicates the quantity of fuel in pounds. The indicator is connected to two fuel transmitters mounted in the fuel cells. One is mounted in the right forward cell and one in the center aft cell. Indicator readings shall be multiplied by 100 to obtain fuel quantity in pounds. Electrical power for operation is supplied from the 115 Vac system and is protected by circuit breaker FUEL QTY in the AC circuit breaker panel. c. Fuel Gage Test Switch. The FUEL GAGE TEST switch (figure 2-6) is used to test the fuel quantity indicator operation. Pressing the switch will cause the indicator pointer to move from the actual reading to a lesser reading. Releasing the switch will cause the pointer to return to the actual reading. The circuit receives power from the 115 Vac system and is protected by a circuit breaker marked FUEL QTY in the AC circuit breaker panel. d. Fuel Pressure Indicator. The fuel pressure indicator (figure 2-6) displays the psi pressure of the fuel being delivered by the boost pumps from the fuel cells to the engine. The circuit receives power from the 28 Vac bus and is protected by the circuit breaker FUEL PRESSURE in the AC circuit breaker panel. e. Fuel Quantity Low Caution Light. The 20 MINUTE FUEL caution light will illuminate when there is approximately 178 pounds of fuel remaining. The illumination of this light does not mean a fixed time period remains before fuel exhaustion, but is an indication that a low fuel condition exists. Electrical power is supplied from the 28 Vdc essential bus. The CAUTION LIGHTS circuit breaker protects the circuit. f. Fuel Boost Pump Caution Lights. The LEFT FUEL BOOST and RIGHT FUEL BOOST caution lights will illuminate when the left/right fuel boost pumps fail to pump fuel. The circuits receive power from the dc essential bus. Circuit protection is provided by the CAUTION LIGHTS, RH FUEL BOOST PUMP and LH FUEL BOOST PUMP circuit breakers.

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2-26. AUXILIARY FUEL SYSTEM. Complete provisions have been made for installing an auxiliary fuel equipment kit in the helicopter cargo/passenger compartment. Two crashworthy bladder type tanks can be installed on the aft bulkhead and transmission support structure allowing the helicopter to be serviced with an additional 300 U.S. gallons (1950 pounds) of fuel. a. Internal Fuel Transfer Switches. Two switches marked INT AUX FUEL LEFT/RIGHT (figure 2-9) are mounted in the ENGINE control panel. Placing

the switches to the forward position energizes the auxiliary fuel system. Fuel is transferred to the main fuels cells. b. Auxiliary Fuel Low Caution Lights. An AUX FUEL LOW caution light is provided to indicate when the auxiliary fuel tanks are empty. The light will illuminate only when the fuel transfer switches are in the forward position, and the auxiliary tanks are empty. The circuit receives power from the 28 Vdc essential bus and is protected by the CAUTION LIGHTS circuit breaker.

Section V. FLIGHT CONTROL SYSTEM

The flight control system is a assisted positive mechanical type, actuated by conventional helicopter controls. Complete controls are provided for both pilot and copilot. The system includes a cyclic system, collective control system, tail rotor system, force trim system and a stabilizer bar.

When an angular displacement of the helicopter/ mast occurs the bar tends to return in its trim plane. The rate at which the bar rotational plane tends to return to a position perpendicular to the mast is controlled by the hydraulic dampers. By adjusting the dampers positive dynamic stability can be achieved, and still allow the pilot complete responsive control of the helicopter.

2-28. CYCLIC CONTROL SYSTEM.

2-29. COLLECTIVE CONTROL SYSTEM.

2-27. DESCRIPTION.

The system is operated by the cyclic control stick (figures 2-4 and 2-7) movement. Moving the stick in any direction will produce a corresponding movement of the helicopter which is a result of a change in the plane of rotation of the main rotor. The pilot cyclic contains the cargo release switch, radio transmitter switch, hoist switch and force trim switch. Desired operating friction can be induced into the control stick by hand tightening the friction adjuster (figure 2-4). a. Synchronized Elevator. The synchronized elevator (figure 2-1) is located on the tailboom. It is connected by control tubes and mechanical linkage to the fore-and-aft cyclic system. Fore-and-aft movement of the cyclic control stick will produce a change in the synchronized elevator attitude. This improves controllability with and increase in CG range. b. Stabilizer Bar. The stabilizer bar is mounted on the main rotor hub trunnion assembly in a parallel plane, above and at 90 degrees to the main rotor blades (figure 2-1). The gyroscopic and inertial effect of the stabilizer bar will produce a damping force in the rotor rotating control system and thus the rotor.

The collective pitch control lever (figure 2-4 and 2-7) controls vertical flight. When the lever is in full down position, the main rotor is at minimum pitch. When the lever is in the full UP position, the main rotor is at maximum pitch. The amount of lever movement determines the angle of attach and lift developed by the main rotor, and results in ascent or descent of the helicopter. Desired operating friction can be induced into the control lever by hand tightening the friction adjuster (figure 2-4). A grip-type throttle and a switch box assembly are located on the upper end of the collective pitch control lever. The pilot switch box contains the starter switch, governor rpm switch, engine idle stop release switch, and landing light/ searchlight switches. A collective lever down lock is located on the floor below the collective lever. The copilot collective lever contains only the grip-type throttle, and governor rpm switch. NOTE The collective pitch control system has a built-in breakaway force to move the stick up from the neutral (center of travel) position of 8 to 10 pounds with boost ON.

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2-30. TAIL ROTOR CONTROL SYSTEM.

2-31. FORCE TRIM SYSTEM.

The system is operated by pilot/copilot anti-torque pedals (figure 2-7). Pushing a pedal will change the pitch of tail rotor blades resulting a directional control. The pedals may also be used to pivot the helicopter about its vertical axis. Pedal adjusters (figure 2-4) are provided to adjust the pedal distance for individual comfort. The copilot directional control pedals are identical to the pilots. A force trim system is connected to the directional controls.

Force centering devices are incorporated in the cyclic controls and directional pedal controls. These devices are installed between the cyclic stick and the h y d ra u l ic s e r v o c y l in d e r s, an d b e t w e e n t h e anti-torque pedals and the hydraulic servo cylinder. The devices furnish a force gradient of “feel” to the cyclic control stick and anti-torque pedals. These forces can be reduced to zero by depressing and holding the force trim pushbutton switch on the cyclic stick grip or by switching the FORCE TRIM switch to the OFF position. A FORCE TRIM ON/OFF switch is installed on the miscellaneous control panel (figure 2-9).

Section VI. HYDRAULIC SYSTEM 2-32. DESCRIPTION. The hydraulic system is used to minimize the force required by the pilot to move the cyclic, collective and pedal controls. A hydraulic pump, mounted on and driven by the transmission supplies pressure to power cylinders. The power cylinders are connected into the mechanical linkage of the helicopter flight control system. Movement of the controls in any direction causes a valve, in the appropriate system, to open and admit hydraulic pressure which actuates the cylinder, thereby reducing the forceload required for control movement. Hydraulic system pressure of 1500 psi is supplied to the cyclic control system. A pressure reducing valve within the system reduces pressure to the collective and directional controls to 1000 psi.

2-33. CONTROL SWITCH. The hydraulic control switch is located on the miscellaneous panel (figure 2-9). The switch is a two-position toggle type labeled HYD CONT ON/ OFF. When the switch is in the ON position, pressure is supplied to the servo system. When switch is in the OFF position the pump outlet is blocked and no pressure is supplied to the system. The switch is fail-safe type. Electrical power is required to turn the switch off.

reservoir and sight gage are visible for inspection through a plastic window in a hole in the transmission fairing.

2-35. HYDRAULIC FILTER. When the hydraulic filter is clogged it will give a visual warning by raising a red indicator button. The red button pops out when the differential pressure across the element exceeds 70 plus or minus 10 psi. Once actuated, the indicator will remain extended until reset manually. When the indicator is in reset position, it will be hidden from view. The indicator is designed not to be actuated by low temperature of fluid, nor by shock loadings, nor by momentary surges in flow. An inspection window is provided to permit ready visual access to the filter indicator. The transparent window is located in the sheet metal structure above the access door on forward face of the cabin island.

2-36. HYDRAULIC PRESSURE CAUTION LIGHT. Low hydraulic system pressure will be indicated by the illumination of HYD PRESSURE segment on the CAUTION panel (figure 2-13). Moderate feedback forces will be noticed in the cyclic when moved.

2-37. ELECTRICAL CIRCUIT. 2-34. RESERVOIR AND SIGHT GLASS. The hydraulic reservoir is a gravity feed type and is located at he right center edge of the roof. The

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Electrical power for hydraulic system is supplied by the 28 Vdc essential bus. The circuit is protected by the HYD CONT circuit breaker.

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Section VII. POWER TRAIN SYSTEMS 2-38. TRANSMISSION.

2-40. INDICATORS AND CAUTION LIGHTS.

The transmission is mounted forward of the engine and coupled to the power turbine shaft at the cool end of the engine by the main driveshaft. The transmission is basically a reduction gearbox, used to transmit engine power at a reduced rpm to the rotor system. A freewheeling unit is incorporated in the transmission to provide a quick-disconnect from the engine if a power failure occurs. This permits the main rotor and tail rotor to rotate in order to accomplish a safe autorotational landing. The tail rotor drive is on the lower aft section of the transmission. Power is transmitted to the tail rotor through a series of driveshafts and gearboxes. The rotor tachometer generator, hydraulic pump, rotor brake disc pad, and dc generator are mounted on and driven by the transmission. A self-contained pressure oil system is incorporated in the transmission. The oil is cooled by an oil cooler and turbine fan. The engine and transmission oil coolers use the same fan. The oil system has a thermal bypass valve which causes the oil to bypass the cooler when the oil is below operating temperature. An oil level sight glass, filler cap, and magnetic chip detector are provided.

a. Transmission Oil Pressure and Temperature Indicator. The transmission oil temperature and pressure indicator is a dual indicator registering temperature (degrees celsius) and pressure (psig) of oil in the transmission. The temperature portion receives temperature indications from an electrical resistance-type bulb and the pressure portion receives pressure indications from the oil pressure transmitter. Electrical power for the circuit is supplied from the 28 Vdc bus and is protected by the XMSN oil press breaker in the DC circuit breaker panel.

2-39. GEARBOXES. a. Intermediate Gearbox - 42-degree. The 42degree gearbox is located at the base of the vertical fin. It provides 42-degree change of direction of the tail rotor driveshaft. The gearbox has a selfcontained wet sump oil system. An oil level sight glass, filler cap, and magnetic chip detector are provided. b. Tail Rotor Gearbox - 90-Degree. The 90-degree gearbox is located at the top of the vertical fin. It provides a 90-degree change of direction and gear reduction of the tail rotor driveshaft. The gearbox has a self-contained wet sump oil system. An oil level sight glass, filler cap, and magnetic chip detector are provided.

2-39.1. DRIVESHAFTS. a. Main Driveshaft. The main driveshaft connects the main output shaft to the transmission input shaft quill. b. Tail Rotor Driveshaft. The tail rotor driveshaft consists of six driveshaft and four hanger bearing assemblies. The assemblies and the 42 degree and 90 degree gearboxes connect the transmission tail rotor drive quill to the tail rotor.

b. Transmission Oil Pressure Low Caution Light. The XMSN OIL PRESS segment in the CAUTION panel (figure 2-13) will illuminate when the transmission oil pressure drops below safe limits. The circuit receives power from the 28 Vdc essential bus. Circuit protection is supplied by the CAUTION LIGHTS circuit breaker. c. Transmission Oil Hot Caution Light. The XMSN OIL HOT segment in the CAUTION panel (figure 2-13) will illuminate when the transmission oil temperature is above safe operating limits. The circuit receives power from the 28 Vdc essential bus and is protected by the CAUTION LIGHTS circuit breaker. d. Transmission and Gearbox Chip Detector. (1) Chip Detector Caution Light. Magnetic inserts are installed in the drain plugs of the transmission sump, 42-degree gearbox, and the 90degree gearbox. When sufficient metal particles collect on the plugs to close the electrical circuit the CHIP DETECTOR segment in the CAUTION panel will illuminate (figure 2-13). A self-closing, springloaded valve in the drain plug permits the magnetic plugs to be removed without the loss of oil. The circuit is powered by 28 Vdc essential bus and protected by the CAUTION LIGHTS circuit breaker. (2) Chip Detector Switch. A CHIP DET switch (figure 2-9) is installed on a pedestal mounted panel. The switch is labeled BOTH, XMSN, and TAIL ROTOR and is spring-loaded to the BOTH position. When the CHIP DETECTOR segment in the CAUTION panel illuminates, reset to extinguish and position switch to XMSN, then TAIL ROTOR, to determine the trouble area. CHIP DET caution light will illuminate when contaminated component is selected. If XMSN is determined to be the trouble area, check the XMSN CHIP IND PANEL located next to pilots feet. Determine from this panel one of three locations as to the trouble area - upper mast, planetary (PLNTY) or sump/oil monitor.

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Section VIII. MAIN AND TAIL ROTOR GROUP 2-41. MAIN ROTOR. a. Description. The main rotor is a semi-rigid, see-saw type. The two all metal blades are connected to a common yoke by blade grips and pitch change bearings with tension straps to carry centrifugal forces. The rotor assembly is connected to the mast with a bearing mounted trunnion joint and secured to the mast with a nut. The nut has provisions for hoisting the helicopter. A stabilizer bar is mounted on the trunnion 90 degrees to the main rotor. Blade pitch change is accomplished by movements of the collective and cyclic control levers, swashplate, and a series of mixing levers and control tubes terminating at the blade grip. The main rotor is driven by the transmission through the mast.

b. RPM Indicator. The rpm indicator is part of the dual tachometer (figure 2-6). The tachometer inner scale displays the rotor rpm. The inner scale pointer is marked with an R.

2-42. TAIL ROTOR. The tail rotor is a two-bladed, semi-rigid delta-hinge type. Each blade is connected to a common yoke by a grip and pitch change bearings. The hub and blade assembly is mounted on the tail rotor shaft with a delta-hinge trunnion to minimize rotor flapping. Blade pitch change is accomplished by movement of the tail rotor pedals which are connected to a pitch control system through the tail rotor (90-degree) gearbox. Blade pitch change serves to offset torque and provide directional control.

Section IX. UTILITY SYSTEM 2-43. PITOT HEATER. The pitot tube (figure 2-1) is equipped with an electrical heater. The PITOT HTR switch is on the left overhead console panel (figure 2-7). ON position actuates the heater in the tube and prevents ice from forming in the pitot tube. Off position deactuates the heater. The electrical circuit for the system receives power from the dc essential bus and is protected by the PITOT TUBE HTR circuit breaker.

2-47. EXTERNAL CARGO REAR VIEW MIRROR. A mirror may be installed under the right lower nose window to give the pilot clear visibility of the external cargo. This mirror may be easily removed and stowed when not in use. A compartment on the right aft side of the forward fuselage, between stations 178 and 211, contains brackets for stowing the external cargo rear view mirror.

2-47.1. WINDSHIELD WIPER. 2-44. HEATED BLANKET RECEPTACLES. Two electrical receptacles are provided to supply 28 Vdc for heated blankets. They are mounted on the inside cabin roof structure aligned with the forward edge of the transmission support structure, one on each side. The electrical circuit for the receptacles receive power from the dc nonessential bus. Circuit protection is provided by the HEATED BLANKET circuit breakers.

2-45. DATA CASE. A data case for maps, flight reports, etc. has been provided and is located at the left side of the pedestal.

2-46. BLACKOUT CURTAINS. Provisions have been made for installing blackout curtains behind pilot and copilot seats and between forward and aft cabin sections. Other blackout curtains may be installed over both cargo door windows and window in removable doorpost. 2-34

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Do not operate the wiper on a dry or dirty windshield. a. Two windshield wipers are provided, one for the right section of the windshield and one for the left section of the windshield. b. The wipers are driven by electric motors with electric power supplied by the dc electrical system. Circuit protection is provided by WINDSHIELD WIPER PILOT and WINDSHIELD WIPER COPILOT circuit breakers on the dc circuit breakers panel. c. The windshield wipers switches on the overhead console mounted MISC panel (Figure 2-7) have five positions: HIGH, MED, LOW, OFF, and PARK. d. The panel also has a selector which permits the operation of windshield wiper for pilot, copilot, or both as desired.

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Section X. HEATING AND VENTILATION 2-48. VENTILATING SYSTEM. a. Description. The ventilating system consists of four independently controlled exterior air scoop ventilators (figure 2-1). Two single orifice air scoops are located on top of the cockpit section and two double orifice air scoops are on top of the cargo/passenger section of the cabin. The amount of air entering the cabin through the ventilators is regulated by the butterfly valve control. b. Operation. Rotate butterfly valve control to desired position to provide outside air for flight.

2-49. HEATING AND DEFROSTING SYSTEM. Two different types of heating and defrosting system may be used on this helicopter. They are the bleed air heater, and the auxiliary exhaust heat exchanger.

Each system is described separately in the following paragraphs. a. Bleed Air Heating and Defrosting System. Aft pedestal outlets are provided. Heat is supplied to all bleed air heaters by the compressor bleed air system. Electrical power for operation of the controls is supplied from the 28 Vdc essential bus and is protected by two circuit breakers marked CABIN HEATER OUTLET VALVE and CABIN HEATER AIR VALVE. Refer to figure 2-10 for controls and their function. b. Auxiliary Exhaust Heater System. The auxiliary exhaust heater system consists of a heat exchanger, and a bleed air driven fan for circulating air through heat exchanger. A mixing valve controls air to maintain the desired temperature. The exhaust heater system controls consist of the cabin heating panel (figure 2-10), a thermostat dial on the right door post and the air directing lever on the pedestal.

Section XI. ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM 2-50. DC AND AC POWER DISTRIBUTION. The dc power (figure 2-11) is supplied by the battery, main generator, standby starter-generator, or the external power receptacle. The 115 Vac power is supplied by the main or standby inverters. The 28 Vac power is supplied by a transformer which is powered by the inverter.

2-51. DC POWER SUPPLY SYSTEM. The dc power supply system is a 28-volt, single conductor system with the negative leads of the generator grounded in the helicopter fuselage structure. In the event of a main generator failure the nonessential bus is automatically deenergized. The pilot may override the automatic action by positioning the NON-ESS BUS switch on the DC POWER control panel to MANUAL ON.

2-52. EXTERNAL POWER RECEPTACLE. The external power receptacle transmits the ground power unit 28 Vdc power to the power distribution system.

NOTE The battery switch shall be in OFF position when GPU is being used. Reverse polarity between helicopter electrical system and GPU can occur. The battery supplies dc power to the power distribution system when the generators and external power receptacle are not in operation.

2-53. MAIN AND STARTER-GENERATOR.

STANDBY

The 28-volt 300 ampere main generator is mounted on and driven by the transmission. This provides generator power during normal flight or autorotations. A standby starter-generator, rated at 300 amperes is mounted on the engine accessory drive section. The standby furnishes generator power in the event of main generator failure.

2-54. DC POWER INDICATORS AND CONTROLS.

NOTE A 7.5 KW (minimum) GPU is required for external starts.

a. Main Generator Switch. The MAIN GEN switch (figure 2-7) is on the overhead console DC

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POWER panel. In the ON position the main generator supplies power to the distribution system and charges the battery. The RESET position is spring-loaded to the OFF position. Momentarily holding the switch to RESET position will reset the main generator. The OFF position isolates the generator from the system. The circuit is protected by the GEN & BUS RESET in the DC circuit breaker panel. b. Battery Switch. The BAT switch (figure 2-7) is located on the DC POWER control panel. ON position permits the battery to supply dc power to the distribution system. ON position also permits the battery to be charged by the generator. The OFF position isolates the battery from the system. c. Starter-Generator Switch. The STARTER GEN switch (figure 2-7) is located on the DC POWER control panel. The START position permits the starter-generator to function as a starter. The STBY GEN position permits the starter-generator to function as a generator and supply dc power to the distribution system. d. Nonessential Bus Switch. The NON-ESS BUS switch (figure 2-7) is located on the DC POWER control panel. The NORMAL ON position permits the nonessential bus to receive dc power from the main generator. The MANUAL ON position permits the nonessential bus to receive dc power from the standby generator when the main generator is off line. e. DC Voltmeter Selector Switch. The VM switch (figure 2-7) is located on the DC POWER control panel. The switch permits monitoring of voltage being delivered from any of the following: BAT, MAIN GEN, STBY GEN, ESS BUS, and NON-ESS BUS. The switch must be rotated to the desired position and voltage is indicated on the dc voltmeter. f. DC Voltmeter. The dc voltmeter (figure 2-6) is located in the center area of the instrument panel and is labeled VOLT DC. Direct Current voltage is indicated on the voltmeter as selected by the VM switch in the overhead console. g. DC Loadmeters - Main and Standby. Two direct current loadmeters are mounted in the lower center area of the instrument panel (figure 2-6). The MAIN GEN loadmeter indicates the percentage of main generator rated capacity being used. The STBY GEN loadmeter indicates the percentage of

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standby generator rated capacity being used. The loadmeters will not indicate percentage when the generators are not operating.

2-55. DC Circuit Breaker Panel. The DC circuit breaker panel (figure 2-11) is located in the overhead console. In the “pushed in” position the circuit breakers provide circuit protection for the 28 Vdc equipment. In the “pulled out” position the circuit breakers deenergize the circuit. In the event of an overload the circuit breaker protecting that circuit will “pop out”. Each breaker is labeled for the particular circuit it protects. Each applicable breaker is listed in the paragraph describing the equipment it protects.

2-56. AC POWER SUPPLY SYSTEM. Alternating current is supplied by two inverters. They receive power from the 28 Vdc essential bus and are controlled from the AC POWER control panel (figure 2-7).

2-57. INVERTERS. Either the main or spare inverter (at the pilots option) will supply the necessary 115 Vac to the distribution system. The inverters also supply 115 Vac to the 28-volt ac transformer which in turn supplies 28 Vac to the necessary equipment. Circuit protection for the inverters is provided by the MAIN INVTR PWR and SPARE INVTR PWR circuit breakers.

2-58. AC POWER INDICATORS AND CONTROLS. a. Inverter Switch. The INVTR switch is located on the AC POWER control panel in the overhead console (figure 2-7). The switch is normally in the MAIN ON position, to energize the main inverter. In the event of a main inverter failure the switch can be positioned to SPARE ON to energize the spare inverter. Electrical power to the INVTR switch is supplied from the dc essential bus. Circuit protection is provided by the INVTR CONT circuit breaker. b. AC Failure Caution Light. The pilot INST INVERTER caution light will illuminate when the inverter in use fails or when the INV switch is in the OFF position.

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c. AC Voltmeter Selector Switch. The AC VM switch is located on the AC POWER control panel (figure 2-7). The switch is used to select any one of the three phases of the 115 Vac three-phase current for monitoring on the ac voltmeter. The three positions on the switch are: AB, AC and BC. Each position indicates that respective phase of the 115 Vac on the ac voltmeter. d. AC Voltmeter. The AC voltmeter is mounted in the center area of the instrument panel (figure 2-6). The ac voltage output from the inverter (main or spare) is indicated on this instrument. The voltage indicated by each of the three selected positions (phases) should be 112 to 118 Vac.

2-59. AC CIRCUIT BREAKER PANEL. The ac circuit breaker panel (figure 2-12) is located on the right side of the pedestal base (figure 2-7). The circuit breakers in the “pushed in” position provide circuit protection for the 28 Vac and 115 Vac operated equipment. The breakers in the “pulled out” position deenergize the circuit. The breakers will “pop out” automatically in the event of a circuit overload. Each breaker is labeled for the particular circuit it protects. Each applicable breaker is listed in the paragraph describing the equipment it protects.

Section XII. LIGHTING 2-60. POSITION LIGHTS. a. General. The position lights consist of eight lights. Two read lights are mounted on the left side of the fuselage, one above and one below the cabin door. Two green lights are mounted on the right side of the fuselage, one above and one below the cabin door. Two white lights are mounted on top of the fuselage, just inboard of the red and green lights. One white light is mounted on the bottom center of the fuselage, and two white lights are mounted on the tail boom vertical fin. Electric power to operate the lights is supplied from the dc essential bus. Circuit protection is provided by NAV LIGHTS circuit breaker in the dc circuit breaker panel. b. Operation. The position lights are controlled from a panel on the overhead console marked EXT LTS (figure 2-7). A three-position switch permits selection of STEADY, OFF or FLASH. Another twoposition switch, controls brilliance and is marked DIM and BRIGHT. When the three-position switch is in STEADY position, all eight navigation lights are illuminated. In FLASH position, the colored lights and the aft white lights flash.

b. Operation. The ON position of the ANTI COLL light switch (figure 2-7) illuminates the anticollision light and starts rotation of the light. OFF position deenergizes the light.

2-62. LANDING LIGHT. a. General. The landing light (figure 2-1) is flushmounted to the underside of the fuselage. It may be extended or retracted to improve forward illumination. Electric power to operate the system is supplied from the dc essential bus. Circuit protection is provided by the LDG LIGHT PWR and LDG & SEARCH LIGHT CONT circuit breakers. b. Operation. The landing light switches (figure 2-4) are on the pilot collective lever switch box. The ON position of the LDG LT ON/OFF switch illuminates the landing light, OFF deenergizes the circuit. The EXT position of the LDG LT EXT/OFF/ RETR extends the landing light to the desired position. RETR position retracts the light. The OFF position stops the light during extension or retraction. The light automatically stops at the full extension/retract position.

2-63. SEARCHLIGHT. 2-61. ANTI-COLLISION LIGHT. a. General. The anti-collision light (figure 2-1) is located on the top aft fuselage area. Electric power to operate the light is supplied from the 28 Vdc essential bus. Circuit protection is provided by the ANTI COLL LIGHT circuit breaker.

a. General. The searchlight is flush-mounted to the underside of the fuselage. The light can be extended and retracted for search illumination. At any desired position in the extend or retract arc, the light may be stopped and rotated to the left or right. Electric power to operate the light is supplied from the 28 Vdc essential bus. Circuit protection is provided by the SEARCHLIGHT PWR and LDG & SEARCHLIGHT CONT circuit breakers.

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b. Operation. (1) Searchlight Switch. The pilot SL switch (figure 2-4) ON position illuminates the light. The OFF position deactivates the light. The STOW position retracts the light into the fuselage well. (2) Searchlight Control Switch. The pilot SEARCH CONT switch (figure 2-4) EXT position extends the light from the fuselage well and moves it forward. RETR position moves the light aft. The L and R position rotates the light left and right.

2-64. DOME LIGHTS. a. General. The dome lights provide overhead lighting for the cabin area. The dome lights are controlled by with a switch on the AFT DOME LTS panel on the roof within easy reach of the cargo passenger station. Electric power to operate the dome light is supplied from the dc essential bus. Circuit protection is provided by the DOME LIGHTS circuit breaker. b. Operation. To operate the dome lights, position the switch to WHITE for white light, RED for red light and OFF to turn lights off. Rotation of the rheostat marked OFF/MED/BRT, increases or decreases the brightness of the dome lights.

2-65. COCKPIT LIGHTS. a. General. Cockpit lights (figure 2-7) are provided for pilot and copilot. One is located on the right of the overhead console and one on the left of the overhead console. The lights receive electric power from the dc essential bus and are protected by the COCKPIT LIGHTS circuit breaker. b. Operation. Rheostat switches are part of each light assembly. Brightness is increased by turning the rheostat clockwise or dimmed by turning counterclockwise. Clockwise rotation of the lens provides white lighting. Counterclockwise rotation of the lens provides red lighting.

2-66. INSTRUMENT LIGHTS. The instrument lights control panel (figure 2-7) is located in the overhead console. The panel contains six switch/rheostats for actuating and controlling the brightness of the various instrument lights. Each switch/rheostat functions the same. OFF position deenergizes the circuit, clockwise rotation increases brightness of the lights, and counterclockwise rotation decreases brightness. The instrument lights all receive electric power from the dc essential bus. They are protected by the CONSOLE PED LIGHTS, INST PANEL LIGHTS and INST SEC LIGHTS circuit breakers.

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a. Pilot Instrument Lights. The pilots instrument lights furnish illumination for the following instruments: gas producer tachometer, torquemeter, engine gas temperature indicator, dual tachometer, airspeed indicator, clock, vertical velocity indicator, turn and slip indicator, altimeter, attitude indicator, radio magnetic indicator, radar altimeter, and standby compass pilot collective switch box. These lights are all on one circuit and are controlled by the switch/rheostat marked PILOT on the INST LTG control panel. Circuit protection is provided by INST PANEL LIGHTS circuit breaker. b. Copilot Instrument Lights. The copilot instrument lights furnish illumination for the instruments on the copilot section of the instrument panel. These instruments consist of an airspeed indicator, attitude indicator, radar altimeter (if installed on UH-1V) altimeter, vertical velocity indicator and radio indicator. The copilot instrument lights are all on one circuit, and are controlled by the switch/rheostat marked COPILOT on the INST LTG console panel. Circuit protection is provided by INST PANEL LIGHTS circuit and copilot 5 VOLT LIGHTS circuit breaker. c. Engine Instrument Lights. The engine instrument lights furnish illumination for the following instruments: transmission oil temperature, fuel quantity, transmission oil temperature/pressure, fuel quantity, engine oil pressure/temperature, loadmeter, AC voltmeter, fuel pressure indicator and DC voltmeter. Each instrument is individually illuminated and control is accomplished by the switch/rheostat marked engine on the INST LTG control panel. Circuit protection is provided by the INST PANEL LIGHTS circuit breaker. d. Secondary Instrument Lights. The four secondary instrument lights are spaced across the top of the instrument panel shield (figure 2-6). These lights furnish secondary illumination for the instrument panel face. The lights are activated and controlled by the switch/rheostat marked SEC on the INST LTG control panel. Circuit protection is provided by the INST SEC LIGHTS circuit breaker.

2-67. OVERHEAD CONSOLE PANEL LIGHTS. The overhead console panel lights (figure 2-7) furnish illumination for all overhead panels.

Each panel is individually illuminated and control is accomplished by the switch/rheostat, marked CONSOLE on the INST LTG control panel. Circuit protection is provided by the CONSOLE PED LIGHTS circuit breaker.

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2-68. PEDESTAL LIGHTS. The pedestal lights (figure 2-7) furnish illumination for the control panels on the pedestal. Each panel is individually illuminated and control is accomplished by the switch/rheostat marked PED on the INST LTG overhead control panel. Circuit protection is provided by the CONSOLE PED LIGHTS circuit breaker.

2-69. TRANSMISSION OIL LEVEL LIGHT. A transmission oil level light is installed to provide illumination to check the transmission oil sight gage. The circuit is actuated by a button-type switch marked XMSN OIL LEVEL LT SWITCH and is located on the right side of the transmission forward bulkhead. Electric power for the transmission oil level light circuit is supplied by the dc battery. Circuit

protection is provided by the battery voltmeter circuit breaker located in the forward radio compartment.

2-70. BAGGAGE COMPARTMENT LIGHTS. The baggage compartment lights system consists of a 28 Vdc circuit breaker, baggage compartment door switch mounted in doorsill, two baggage compartment lights and a PRESS-TO-TEST (BAGGAGE FIRE/DOOR TEST) switch. NOTE Part of door switch serves to energize instrument panel warning light in case baggage compartment door is open. If the light illuminates, land as soon as possible.

Section XIII. FLIGHT INSTRUMENTS 2-71. AIRSPEED INDICATORS. The pilot and copilot airspeed indicators (figure 2-6) display the helicopter indicated airspeed (IAS) in knots. The IAS is obtained by measuring the difference between impact air pressure from the pitot tube and the static air pressure from the static ports located in the pitot tube. IAS is inaccurate due to instrument and installation errors. Refer to Chapter 7 for the airspeed calibration correction chart. NOTE IAS below 20 KIAS is inaccurate due to rotor downwash.

2-72. TURN AND SLIP INDICATOR. The turn and slip indicator (4 MIN TURN) (figure 26) displays the helicopter slip condition, direction of turn and rate of turn. The ball displays the slip condition. The pointer displays the direction and rate of the turn. The circuit receives power from the dc essential bus and is protected by the TURN & SLIP IND circuit breaker.

2-73. VERTICAL VELOCITY INDICATOR. The vertical velocity indicator (figure 2-6) displays the helicopter ascent and descent speed in feet per

minute. The indicator is actuated by the rate of atmospheric pressure change.

2-74. ALTIMETER INDICATOR. The altimeter (ALT) (figure 2-6) furnishes direct readings of height above sea level and is actuated by the pitot static system. Two altimeters are provided, one for the pilot and one for the copilot.

2-75. ATTITUDE INDICATORS. a. Pilot Attitude Indicator. The pilot attitude indicator is located on the pilot section of the instrument panel (figure 2-6). The indicator displays the pitch and roll attitude of the helicopter. An OFF warning flag in the indicator is exposed when electrical power to the system is removed. However, the OFF flag will not indicate internal system failure. The attitude indicator has an electrical trim in the roll axis in addition to the standard pitch trim. Degrees of pitch and roll are indicated by a universally mounted sphere. The horizon is represented as a white bar on the sphere; horizontal markings indicate the degree of drive or climb, while bank (roll) angles are read from the semicircular scale located on the upper half of the indicator face. The pitch trim knob, located on the lower right corner of the indicator, is adjusted to center the horizon on the indicator sphere with the center mark

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on the roll (bank) indicator in regard to the normal flight attitude of the helicopter. The attitude indicator is operated by 115 Vac power, supplied by the inverter. Circuit protection is provided by the PILOT ATTD circuit breakers in the AC circuit breaker panel.

Caging of the copilot (J-8) gyro attitude indicator should be kept to a minimum, and should be caged in flight only when the helicopter is in straight and level flight. The caging knob shall never be pulled violently. b. Copilot Attitude Indicator. The copilot attitude indicator (figure 2-6) is located in the copilot section of the instrument panel. It is operated by 115 Vac power supplied by the inverter. Circuit protection is provided by the COPILOT ATTD circuit breakers in the AC circuit breaker panel. In a climb or dive exceeding 27 degrees of pitch the horizontal bar will stop at the top or bottom of the case and the sphere then becomes the reference. The copilot attitude indicator must be caged manually by pulling the PULL TO CAGE knobs smoothly away from the face of the instrument to the limit of its travel and then releasing quickly.

2-76. FREE-AIR TEMPERATURE INDICATOR (FAT). The free-air temperature indicator (figure 2-7) is located at the top center area of the windshield. The indicator displays the outside air temperature in degrees celsius.

2-77. STANDBY COMPASS. The standby (magnetic) compass (figure 2-6) is mounted in a bracket at the center right edge of the instrument panel. A compass correction card is located in the card holder above the compass.

2-78. FIRE DETECTOR WARNING SYSTEM. a. A FIRE detector warning light (figure 2-6) is located in the upper right section of the instrument panel. The press-to-test (FIRE DETECTOR TEST)

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test switch is located to the left of the fire warning light. Excessive heat in the engine compartment causes the FIRE light to illuminate. Depressing the press-to-test switch also causes the light to illuminate for testing. Electrical power for the circuit is supplied from the 28 Vdc essential bus and is protected by the FIRE DET circuit breaker. b. A BAGGAGE FIRE detector warning light is located on the instrument panel. The press-to-test (BAGGAGE FIRE/DOOR TEST) switch is located near the baggage fire warning light. This is a smoke detector which is a closed assembly, solid state, electronic component and a light sensitive detector. The smoke detector, located in forward end of baggage compartment roof, is protected from baggage by a protective guard. When smoke reduces light transmission in baggage compartment 30 to 35 percent below that of clean air, the smoke detector will send a signal to fire detector control amplifier. This will cause fire detector light on instrument panel to illuminate (flashing ON and OFF) intermittently. Electric power for this circuit is supplied from the 28 Vdc bus and is protected by the BAGGAGE SMOKE circuit breaker. NOTE Press-To-Test (BAGGAGE FIRE/DOOR TEST) switch will test the ENG OIL FILTER warning light.

2-79. MASTER CAUTION SYSTEM. a. Master Caution Indicator. The master caution indicator light (figure 2-6) on the instrument panel will illuminate when fault conditions occur. This illumination alerts the pilot and copilot to check the caution panel for the specific fault condition. b. Caution Panel. The CAUTION panel (figure 213) is located on the pilot side of the pedestal. Worded segments illuminate to identify specific fault conditions. The worded segments are readable only when the light illuminates. When a light illuminates, indicating a fault condition, it will remain illuminated until the fault condition is corrected. Refer to figure 2-13 for explanation of the fault condition. (1) Bright-Dim Switch. The BRIGHT-DIM switch on the CAUTION panel permits the pilot to manually select a bright or dimmed condition for all the individual worded segments and the master caution indicator. The dimming switch position will

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work only when the pilot instrument lights are on. The master caution system lights will be in bright illumination after each initial application of electrical power; when the pilot instrument lights are turned OFF, or a loss of the power from the dc essential bus occurs. (2) Reset-Test Switch. The RESET-TEST switch on the CAUTION panel enables the pilot to manually reset and test the master caution system. Momentarily placing the switch in the RESET position, extinguishes and resets the master caution indicator light so it will again illuminate should another fault condition occur. Momentarily placing the switch in TEST position will cause the illumination of all the individually worded segments and the master caution indicator. Only the lamp circuitry is tested; the condition circuitry is not. Testing of the system will not change any particular combination of fault indications which might exist prior to testing.

NOTE The worded segments will remain illuminated as long as fault condition or conditions exist, unless the segment is rotated. c. Electric Power. Electric power for the master caution system is supplied from the 28 Vdc essential bus. Circuit protection is provided by the CAUTION LIGHTS circuit breakers.

2-80. RPM HIGH — LOW LIMIT WARNING SYSTEM. The system provides an immediate warning to check the dual tachometer for high or low rotor rpm and/or engine rpm. The rpm light warning and audio

wa r n i n g f u n c t i o n s w h e n t h e f o l l o w i n g r p m conditions exist: Warning light only

For rotor rpm of 324 (plus or minus 5) or engine rpm of 6800 (plus or minus 100). (High Warning)

Warning light and audio warning signal in combination

For rotor rpm of 294 engine rpm of 6100 (plus or minus 100) or both. (Low Warning) NOTE

The audio warning signal will be heard in the headsets. The signal is a varying oscillating frequency, starting low and building up to a high pitch. Signal alternates on for 0.85 second, then off for 1 second. a. RPM Warning Light. The high-low warning light is located on the upper left area of the pilot instrument panel (figure 2-6). The light illuminates to provide a visual warning of high or low rotor rpm and/or engine rpm. b. RPM Switch — Low RPM Audio. The LOW RPM AUDIO/OFF switch is on the engine control (figure 2-9). When in the OFF position, the switch prevents audio warning signal from functioning during engine starting and shutdown and with electrical power on without the engine running. This eliminates use of the circuit breaker as a switch, and increases safety by having warning light working at all times. The LOW RPM AUDIO/ OFF switch is spring loaded to the AUDIO position. When the switch has been manually turned to OFF for engine starting, it will automatically return to the AUDIO position when normal operating range is reached.

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Section XIV. SERVICING, PARKING AND MOORING 2-81. SERVICING. (FIGURE 2-14.) 2-82. APPROVED COMMERCIAL FUELS, OILS, AND FLUIDS. (FIGURE 2-14) 2-82.1. FUEL SYSTEM SERVICING.

(c) Activate flow control handle to ON or FLOW position. Fuel flow will automatically shut off when fuel cell is full. Just prior to normal shut off fuel flow may cycle several times as maximum fuel level is reached. (d) Assure that fuel control handle is in OFF or NO FLOW position and remove nozzle. (7) Gravity or open port: (a) Remove fuel filler cap.

Servicing personnel shall comply with all safety precautions and procedures specified in FM 10-68 Aircraft Refueling Field Manual.

(b) Using latch tool attached to filler cap cable open refueling module. (c) Remove nozzle cap and insert nozzle into fuel receptacle. (d) Fill to specified level.

a. Refer to figure 2-14 for fuel capacities. b. Refer to figure 2-14 for approved fuel. c. The helicopter is serviced as follows: (1) Refer to figure 2-14 for fuel fillers locations.

(e)

Remove nozzle.

(f) Close refueling module by pulling cable until latch is in locked position. Refer to figure 2-14. (8) Replace fuel nozzle cap. (9) Replace fuel filler cap. (10)Disconnect fuel nozzle ground.

(2) Assure that fire guard is in position with fire extinguisher.

(11)Disconnect servicing unit.

ground

from

helicopter

to

(3) Ground servicing unit to ground stake.

(12)Disconnect servicing unit ground from ground stake.

(4) Ground servicing unit to helicopter.

(13)Return fire extinguisher to designated location.

(5) Ground fuel nozzle to ground receptacle located adjacent to fuel receptacle on helicopter.

d. Rapid (Hot) Refueling. (1) Before rapid refueling. (a) Throttle - idle. (b) FORCE frictioned.

Ensure that servicing unit pressure is not above 125 psi while servicing.

(c)

TRIM

-

ON

or

controls

Refuel as describe in paragraph c.

above.

(a) Remove fuel filler cap and assure that refueling module is in locked position. Refer to figure 2-14.

(2) During rapid refueling. A crewmember should observe the refueling operation (preformed by authorized refueling personnel) and stand fire guard as required. One crewmember shall remain in the helicopter to monitor controls. Only emergency radio transmission should be made during RAPID refueling.

(b) Remove nozzle cap and insert nozzle into fuel receptacle and lock into place.

(3) After rapid refueling, the pilot shall be advised by the refueling crew that fuel cap is secure and grounding cables have been removed.

(6) Closed circuit.

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2-83. TYPES AND USE OF FUELS. a. Fuel Types. (1) Army Standards Fuels. These are the Army-designated primary fuels adopted for worldwide use, and are the only fuels available in the Army supply system. (2) Alternate Fuels. These are fuels which can be used continuously when Army standard fuel is not available, without reduction of power output. Power setting adjustments and increased maintenance may be required when an alternate fuel is used. (3) Emergency Fuels. These are fuels which can be used if Army standard and alternate fuels are not available. Their use is subject to a specific time limit. b. Use of Fuels (1) There is no special limitation on the use of Army standard fuel, but certain limitations are imposed when alternate or emergency fuels are used. For the purpose of recording, fuel mixtures shall be identified as to the major component of the mixture except when the mixture contains leaded gasoline. A fuel mixture which contains over 10 percent leading gasoline shall be recorded as all leaded gasoline. The use of any fuels other than standard will be recorded in the FAULTS/ REMARKS column of DA Form 2408-13, Aircraft Maintenance and Inspection Record, noting the type of fuel, additives, and duration of operation. (2) The use of kerosene fuels (JP-5 type) in turbine engines dictates the need for observance of special precautions. Both ground starts and air restarts at low temperature may be more difficult due to low vapor pressure. Kerosene fuels having a freezing point of -40 degrees F (-40 degrees C) limit the maximum altitude of a mission to 28,000 feet under standard day conditions. Those having a

freezing point of -67 degrees F (-53 degrees C) limit the maximum altitude of a mission to 33,000 feet under standard day conditions. (3) Mixing of Fuel in Helicopter Tanks. When changing from one type of authorized fuel to another, for example JP-4 to JP-5, it is not necessary to drain the helicopter fuel system before adding the new fuel. (4) Fuels having the same NATO code number are interchangeable. Jet fuels conforming to ASTMD-1655 specification may be used when MILT-5624 fuels are not available. This usually occurs during cross country flights where helicopters using NATO F-44 (JP-5) are refueled with NATO F-40 (JP-4) or commercial ASTM type B fuels. Whenever this condition occurs, the engine operating characteristics may change in that lower operating temperature, slower acceleration, lower engine speed, easier starting, and shorter range may be experienced. The reverse is true when changing from F-40 (JP-4) fuel to F-44 (JP-5) or commercial ASTM type A-1 fuels. Specific gravity adjustments in fuel controls and flow dividers shall be set for the type of fuel used. Most commercial turbine engines will operate satisfactorily on either kerosene or JP-4 type fuel. However, the difference in specific gravity may possibly require fuel control adjustments; if so, recommendations of the manufacturers of the engine and airframe are to be followed.

2-84. GROUND HANDLING EQUIPMENT, COVERS, ROTOR TIEDOWNS, AND MOORING DIAGRAM. (FIGURE 2-15.)

NOTE All tiedowns shall be tight.

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Figure 2-15. Ground handling equipment, covers, rotor tiedowns and mooring diagram

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Section XV. AUXILIARY EQUIPMENT 2-85. MAIN ROTOR BRAKE (IF INSTALLED). a. General. A rotor brake is provided for stopping rotation of the main rotor blade and to prevent the rotor blade from turning while the helicopter is parked. The rotor brake handle is located overhead within reach of the pilot (figure 2-7). The ROTOR BRAKE warning light, with a built-in press to test, is located on the instrument panel above the dual tachometer (figure 2-6). Electric power for the light is supplied from the 28 Vdc bus. Circuit protection is provided by the ROTOR BRAKE WARNING LIGHT circuit breaker (figure 2-11). The reservoir capacity and fluid specification are shown on the Servicing Diagram (figure 2-14). b. Testing — Rotor Brake Warning System. Test the rotor brake warning system, before engine start, by pressing the press to test ROTOR BRAKE warning light. The rotor brake warning system can also be tested by pulling on the rotor brake handle

enough to disengage the handle from the detent position. c. Operation — Rotor Brake. NOTE Do not apply rotor brake above 130 rotor RPM except in emergencies. (1) Pull down on the rotor brake handle toward vertical position. For maximum braking, do not move handle past center. (2) To release brake, slowly return handle to horizontal position into the detent position. (3) For parking the helicopter, pull the brake handle as for maximum braking effort but continue on past the center position (handle past vertical) to parking position. While in parking position the brake does not exert maximum braking effort.

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CHAPTER 3 AVIONICS Section I. COMMUNICATIONS 3-1.

GENERAL.

This chapter covers the electronic equipment configuration in UH-1H II helicopters. It includes a brief description and technical characteristics of the electronic equipment, capabilities, and location. This chapter also contains complete operating instructions for all signal equipment. For mission avionics equipment, refer to Chapter 4, Mission Equipment.

3-2. ELECTRONIC EQUIPMENT CONFIGURATION. The configuration consists of headset cordage, keying switches, and the equipment listed in figure 3-1. a. Headset Cordage. The pilot and copilot cordage connectors are located at their respective sides near the aft portion of the overhead console. The crew cordage connectors are located near the overhead mounted signal distribution panel (ICS) at each crew station. b. Keying Switches. A trigger type keying switch is located on each (pilot and copilot) cyclic control stick grip. The half depressed position of the trigger switch is used for keying the interphone. The fully depressed (second detent) position of the trigger switch keys the radio selected with the transmitinterphone selector switch on the signal distribution panel. A foot-operated type keying switch (pilot and copilot) is located at each side of the center console, between the center console and cyclic control stick, and on the cabin floor at each crew station. The depressed position of the foot-operated switch keys the radio or interphone selected with the rotary selector switch at the appropriate signal distribution panel. c. Power Supply and Circuit Breakers. Refer to figure 3-2.

3-3. SIGNAL DISTRIBUTION PANEL (ICS C-6533/ARC). a. Description. Two panels are installed in the pedestal for the pilot and copilot and two panels are installed in the cabin roof aft of the overhead console for the right and left crewmembers. The system is used for intercommunications and radio control. The system has three modes of operation; two way radio communication, radio monitoring, and interphone. b. Controls and Functions. Refer to figure 3-3. c. Operating Procedures. (1) NAV receiver switch — As desired. (2) AUX receiver switch — As desired. (3) Transmit-interphone selector switch — As desired. (4) Receiver switches — As desired. (5) HOT MIKE switch — As desired. (6) VOL control — Adjust.

3-4.

UHF RADIO SET AN/ARC-159. NOTE Helicopter has provisions for AN/ARC159 UHF radio set.

a. Description. The radio set provides two way communications in the UHF (225.00 to 399.9mHz) band, located at the left side of the pedestal. The transmitter and receiver operate on the same frequency and are simultaneously tuned by frequency selector switch is positioned to BOTH the main and guard receivers are both energized. This provides constant monitoring of the main receivertransmitter frequency setting.

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Figure 3-1. Electronic equipment configuration

3-2

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Figure 3-2. Power supply and circuit breakers

3-3

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Figure 3-3. Signal distribution panel (ICS)

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(4) AUDIO — Adjust.

b. Controls and Functions. Refer to figure 3-4. c. Operating Procedures. (1) UHF Mode selector — As desired. (2) VOL — Adjust. (3) SQUELCH — As desired. (4) Function selector — As desired. (5) Frequency — Select. (6) ICS Transmit-interphone selector switch — No. 2 position. (7) ICS Receiver switch No. 2 — ON. (8) To transmit — Depress foot switch or cyclic stick trigger switch to the second detent and speak into microphone. d. Emergency Operation. (1) Mode Selector — Guard. (2) Function Selector — Both.

3-5.

VHF RADIO SET — AN/ARC-115. NOTE Helicopter has provisions for AN/ARC115 VHF radio set.

a. Description. The VHF Radio Set provides amplitude-modulated, narrow band voice communications within the frequency range of 116.000 to 149.975 mHz on 1360 channels for a distance of approximately 50 miles line of sight. A guard receiver is incorporated and fixed tuned to 121.50 mHz. The panel is labeled VHF AM COMM and mounted on the left side of the pedestal.

(5) Transmit-interphone selector — No. 3 position. (6) To transmit — Depress foot switch or cyclic stick trigger switch to the second detent and speak into the microphone. d. Emergency Operation. Function selector — T/ R GUARD.

3-6.

FM RADIO SET — AN/ARC-114A. NOTE The ARC-114A or ARC-114 may be installed in the helicopter. They are basically the same. Refer to figure 3-6 for the differences.

a. Description. The FM Radio Set provides two way frequency modulated (FM) narrow band voice communications and homing capability within the frequency range of 30.00 to 75.95 mHz on 920 channels for a distance range limited to line of sight. A guard receiver is incorporated in the set and is fix tuned to 40.50 mHz. The set is marked VHF FM COMM and is mounted on the center console. b. Controls and Functions. Refer to figure 3-6. c. Operating Procedures. d. Two way voice communication. (1) Function selector — As desired. (2) Frequency — Select. (3) RCVR TEST — Press to test. (4) AUDIO — Adjust. (5) Receiver switch No. 5 — ON.

b. Controls and Functions. Refer to figure 3-5. c. Operating Procedures. (1) Function selector — As desired.

(6) Transmit-interphone selector — No. 5 position. (7) To transmit — Depress foot switch or cyclic stick trigger switch to the second detent and speak into the microphone.

(2) Frequency — Select. (3) RCVR TEST — Press to test.

e. Emergency Operation. Function selector — T/R GUARD. Select 40.50 mHz.

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Figure 3-4. UHF AM AN/ARC-159 control panel (Sheet 1 of 2)

3-6

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Figure 3-4. UHF AM AN/ARC-159 control panel (Sheet 2 of 2)

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Figure 3-5. Control panel, AN/ARC-115

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Figure 3-6. Control panel AN/ARC-114 (Sheet 1 of 2)

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Figure 3-6. Control panel AN/ARC-114A (Sheet 2 of 2)

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3-7.

RADIO SET AN/ARC-102.

b. Controls and Functions. Refer to figure 3-7.

a. The AN/ARC-102 is a long range High Frequency (HF) Single Side Band (SSB) transceiver which transmits and receives in the 2.0 to 30 megahertz frequency range. The set tunes in one kHz stops to any one of 28,000 manually selected frequencies. The primary mode of operation is SSB, however, the ARC-102 can also transmit and receive a compatible AM signal. The receiver-transmitter is composed of eleven plug-in modules, which includes an interchangeable internal power supply. The complete unit is contained in a metal case and weighs 50 pounds. The ARC-102 antenna coupler is mounted in the forward section of the tail boom. The coupler automatically matches the impedance of the long wire antenna to the channel frequency selected on the remote control unit. Power to operate the antenna coupler is supplied from the receivertransmitter.

c. Operating Procedures. (1) Function Selector — As desired. (2) Frequency — Select. (3) Transmit-interphone selector — No. 4 position. (4) Receiver Switch No. 4 — ON. (5) RF SENS comfortable level.

Control

—

Adjust

to

a

(6) To transmit — Depress foot switch or cyclic stick trigger switch to the second detent and speak into the microphone.

Section II. NAVIGATION (2) Manual Operation.

NOTE Helicopter 83ADF.

3-8.

has

provisions

for

ARN-

(a) Mode selector switch — LOOP.

DIRECTION FINDER SET — AN/ARN-83.

a. Description. The Direction Finder set provides radio aid to navigation within the 190 to 1750 kHz frequency range. In automatic operation, the set presents continuous bearing information on pointer No. 1 of RMI, to any selected radio station. The set simultaneously provides aural reception of the station's transmission. In manual operation, the operator determines the bearing to any selected radio station by controlling the aural null of the directional antenna. The set may also be operated as a receiver. b. Controls and Functions. Refer to figure 3-8. c. Operating Procedures. (1) Automatic Operation. (a) ICS receiver switch (NAV) — ON. (b) Mode selector switch — ADF. (c)

Frequency — Select.

(d) Volume — Adjust.

(b) BFO switch — ON. (c) LOOP L /R switch — Press right or left and rotate loop for null.

3-9.

DIRECTION FINDER SET — AN/ARN-89.

a. Description. The Direction Finder set provides radio aid to navigation within the 100 to 3000 kHz frequency range. The set has three operating modes selected by the function switch on the ADF control panel. When operating in the COMP mode the set presents continuous bearing information on pointer No. 1 of the RMI, to the selected radio station. The ANT mode permits reception of radio range navigation or radio broadcast stations. The LOOP mode is used for aural null homing and manual direction finder. A beat frequency oscillator is included to provide an audible indication for identification and tuning. It is energized by the CW switch. ADF bearing signals are presented on the pilot and copilot course indicators. b. Controls and Functions. Refer to figure 3-9.

Rev. 1

3-11

BHT PUB-92-004-10

Figure 3-7. HF Radio control panel

3-12

BHT PUB-92-004-10

Figure 3-8. Direction finder control panel C6899/ARN-83 c. Operating Procedures. (1) Automatic Operation. (a) ICS receiver switch (NAV) — ON. (b) Mode selector switch — COMP. (c)

Frequency — Select.

(d) Volume — Adjust. (2) Manual Operation. (a) LOOP L/R switch — Position to the right or left and rotate a loop for null. (b) CW/Voice/Test switch — CW. (c) LOOP L /R switch — Position to the right or left and rotate loop for null.

3-10. RADIO RECEIVING SET — AN/ARN-82. a. The AN/ARN-82 Navigation Receiver provides for reception of 200 channels with 50 kHz spacing.

This permits reception of VHF visual omni-direction range (VOR) signals and localizer signals. Localizer frequencies are all the odd tenth mHz frequencies between 108.00 mHz and 112.0 mHz. The localizer function is energized when these frequencies are selected. Localizer and VOR signals are presented aurally through the intercom system. Localizer signals are also presented visually via the vertical pointer of the course indicator, and VOR signals are presented visually via the course indicator and the No. 2 pointer of the bearing heading indicator. b. Controls and Functions. Refer to figure 3-10. c. Operating Procedures. Refer to figure 3-11. (1) Battery switch — ON (OFF for APU). (2) VHF NAV OMNI and INTERCOM circuit breakers — IN. (3) Power switch — PWR. (4) Frequency — Select.

Rev. 1

3-13

BHT PUB-92-004-10

Figure 3-9. Direction finder control panel C-7392/ARN-89

3-14

BHT PUB-92-004-10

Figure 3-10. Navigation control panel C6873/ARN-82 (5) AUX switch on C-6533/ARC panel — UP. (6) AN/ARN-82 Stopping Procedure. Power switch — OFF.

3-11. GYROMAGNETIC COMPASS SET. a. Description. (1) The gyromagnetic compass set is a direction sensing system which provides a visual indication of the magnetic heading (MAG) of the helicopter. The information which the system supplies may be used for navigation and to control flight path of the helicopter. The system may also be used as a free gyro (DG) in areas where the magnetic reference is unreliable. (2) A radio magnetic indicator is installed in the pilot instrument panel. A second radio magnetic indicator (RMI) is installed in the copilot instrument panel. The copilot indicator is a repeater type instrument similar to the pilot indicator except that it has no control knobs. The moving compass card on both indicators displays the gyromagnetic compass heading. The number 1 pointer on the

indicators indicates the bearing to the station selected on the ADF receiver. The number 2 pointer indicates VOR bearing. b. Controls and Functions. Refer to figure 3-12. c. Operating Procedures. (1) INV switch — MAIN and STBY. (2) Radio magnetic indicator (pilot only) — Check power failure indicator is not in view. (a) Slaved Gyro Mode (figure 3-12). 1.

COMPASS switch — MAG.

2. Synchronizing (Null) annunciator.

knob

—

Center

NOTE The system does not have a "fast-slewing" feature. If the compass is 180 degrees off the correct helicopter heading when the system is energized it will take approximately 30 minutes for the compass to slave to the correct headings.

Rev. 1

3-15

BHT PUB-92-004-10

Figure 3-11. Course selector indicator ID 1347/ARN-82

3-16

BHT PUB-92-004-10

Figure 3-12. Gyromagnetic compass indicator (RMI)

3-17

BHT PUB-92-004-10

3.

Magnetic heading — Check.

(b) Free Gyro Mode. 1.

COMPASS switch — DG.

2.

Synchronizing knob — Set heading.

3. Annunciator — Center position and then does not change (annunciator is de-energized in the free gyro (DG) mode). (c)

Inflight Operation.

1. Set the COMPASS switch to DG or MAG as desired for magnetically slaved or free gyro mode of operation. Free gyro (DG) mode is recommended when flying in latitudes higher than 70°. 2. When operated in the slaved (MAG) mode, the system will remain synchronized during normal flight maneuvers. During violent maneuvers the system may become unsynchronized, as indicated by the annunciator moving off center. The system will slowly remove all errors in synchronization; however, if fast synchronization is desired turn the synchronizing knob in the direction indicated by the annunciator until the annunciator is centered again.

a. Description. The Marker Beacon Receiver set is a radio aid to navigation. It receives 75 mHz marker beacon signals from a ground transmitter to provide the pilot with aural and visual information. The marker beacon controls and indicator are located on the instrument panel to aid in determining helicopter position for navigation or instrument approach. b. Control and Functions. Refer to figure 3-13.

3-13. RADAR ALTIMETER — GENERAL. a. General. The radar altimeter, located in the lower right center of the pilots instrument panel, is an electronic low-level altitude system which provides the following features (figure 3-14): (1) The radar altimeter provides precise altitude indications from 0 to 1500 feet. The indicator contains a test button and a DH set knob control. (2) The system includes a manually set, low level warning lamp to warn that a preselected low altitude limit has been reached. (3) The indicator contains an OFF flag that indicates the system is not receiving power. b. Operation — Radar Altimeter.

(3) When operating in the free gyro (DG) mode, periodically update the heading to a known reference by rotating the synchronizing knob.

(1) Radar altimeter circuit breaker — CLOSED OFF flag disappears after approximately 35 seconds and pointer indicates 0 ± 5 feet.

3-12. MARKER BEACON RECEIVER.

(2) Select desired DH warning index by rotating control knob until indicator index pointer reaches desired ground clearance setting.

NOTE Helicopter has provisions for Marker Beacon Receiver.

3-18

Rev. 1

(3) Test system by depressing TEST button. Altitude pointer should indicate 100 ± 20 feet. Release TEST button.

BHT PUB-92-004-10

Figure 3-13. Marker beacon controls

Figure 3-14. Radar altimeter

3-19

BHT PUB-92-004-10

Section III. TRANSPONDER AND RADAR 3-14. TRANSPONDER SET AN/APX-72. a. Description. The APX-72 provides radar identification capability. Five independent coding modes are available to the pilot. The first three modes may be used independently or in combination. Mode 1 provides 32 possible code combinations, any one of which may be selected in flight. Mode 2 provides 4,096 possible code combinations but only one is available since the selection dial is not available in flight. Mode 3/A provides 4,096 possible codes, any of which may be selected in flight. Mode C in this installation is not utilized. Mode 4, which is connected to an external computer, can be programmed prior to flight to display any one of many classified operational codes for security identification. The effective range depends on the capability of interrogation radar and line-of-light. b. Controls and Functions. Refer to figure 3-15. c. Starting Procedure. (1) Preliminary. (a) MASTER control — OFF. (b) INDENT-MIC switch — OUT. (c) M-1, M-2, M-3/A, M-C and MODE 4 switches — OUT. (d) AUDIO-LIGHT switch — OUT. (e)

RAD TEST-MON switch — OUT.

(f) MODE 1, 3/A and 4 code select switches — Set to operational code required. (g) MODE 2 code select switch — Set to operational code required. (2) Starting.

LOW — low receiver sensitivity for receiving high energy signals. NORM — normal receiver sensitivity. EMER — refer to step f. (b) M-1, M-2, M-3/A and MODE 4 switches ON as required. (c)

AUDIO-LIGHT switch — LIGHT.

(d) INDENT-MIC switch — OUT. (e)

RAD TEST-MON switch — MON.

d. Normal Operation. (1) MASTER control — LOW or NORM as required. (2) M-1, M-2, M-3/A, and MODE 4 switches ON — unless operational requirements indicate that only specific modes are to be used, then all other mode switches will be OUT. (3) AUDIO-LIGHT switches — LIGHT. (4) INDENT-MIC switch — OUT. (5) RAD TEST-MON switch — OUT. e. Identification of Position (I/P) Operation. The APX-72 will transit position identifying signals to all interrogating stations on modes 1, 2, and 3/A when the INDENT-MIC switches on the control panel is set the IDENT. Transmission of the I/P signal will occur in these modes even if the mode enable switches are in the OUT position. the I/P operation is as follows: Momentarily hold the INDENT-MIC switch in the IDENT position (spring-loaded return) and then release it. This action will cause the APX-72 to transit the I/P signal for a period of approximately 25 seconds to all interrogating stations on modes 1, 2, and 3/A. Repeat as required.

(a) MASTER control. STBY — one minute for standard temperature conditions and two minutes under extreme ranges of operating temperature.

3-20

f. Emergency Operation. During an aircraft emergency or distress condition, the APX-72 may be used to transit specially coded emergency signals on modes 1, 2, and 3/A to all interrogating stations.

BHT PUB-92-004-10

Figure 3-15. Transponder set AN/APX-72

3-21

BHT PUB-92-004-10

The emergency signals will be transmitted as long as the MASTER control on the control panel remains in the EMER position. For emergency operation, set the controls as follows:

(1) MASTER control — EMER — leave in that position for the duration of the emergency.

(2) MASTER control — NORM or LOW after emergency is over.

AUDIO LIGHT switch on the control panel as follows: (1) AUDIO-LIGHT switch — AUDIO — Mode 4 interrogating and reply pulses will be audible in the pilots headset and visible on the RELAY light. (2) AUDIO-LIGHT switch — LIGHT — Indication of mode 4 interrogating and reply pulses will be visible on the REPLY light. h. Stopping Procedure. (1) MASTER control — OFF.

g. Monitoring Operation. Valid mode 4 Interrogations and replies can be monitored either aurally and visually or visually by placement of the

3-22

(2) IDENTI-MIC switch — OUT. (3) M-1, M-2, M-3/A, and MODE 4 switches OUT.

BHT PUB-92-004-10

CHAPTER 4 MISSION EQUIPMENT 4-1.

(c) CARGO RELEASE switch ó ARM; check CARGO RELEASE ARMED light illuminates (Figure 4-1A).

CARGO HOOK.

Helicopters equipped with a non-rotating cargo suspension unit, which maintains the hook in a fixed position, (facing forward) should be used only with a cargo sling having a swivel attachment ring. a device which may be used for this application is: Sling, Endless, Nylon Webbing, Type I 10 inch, Part No. PD 101-10. a. Description. External cargo can be carried by means of a short single cable suspension unit (figure 4-1), secured to the primary structure and located at the approximate center of gravity. This method of attachment and location has proved to be the most satisfactory for carrying external cargo. Pitching and rolling due to cargo swinging is minimized, and good stability and control characteristics are maintained under load. A MANUAL CARGO RELEASE PUSH pedal is located between the pilot tail rotor control pedals, and an electrical release pushbutton switch is on the cyclic control stick. Before the electrical release switch on the cyclic control stick can be actuated, the CARGO REL switch on the overhead panel must be positioned to ARM. When not in use the cargo suspension unit need not be removed, nor does it required stowing. Three cable and spring attachments keep the unit centralized, and the hook protrudes only slightly below the lower surface of the helicopter. A rear view mirror enables the pilot to visually check operation of the external cargo suspension hook. b. Preflight Procedures. For external load class and limit, refer to figure 4-1A. (1) Check condition and security of cargo suspension unit. (2) Check operation of cargo release as follows: HOOK

NOTE Pedal release will function regardless of CARGO REL switch position. (e) Cargo release pedal ó Push and hold; pull down on cargo hook, hook should open. Release pedal and cargo hook; hook should close and lock (f) Check primary load ring and secondary load ring for condition and proper size (Table 4-1). Check for correct rigging.

USE OF INAPPROPRIATELY SIZED LOAD RINGS MAY RESULT IN LOAD HANG-UP WHEN LOAD RING IS T OO S MA LL OR IN ADV ERTEN T LOAD RELEASE IF LOAD RING IS TOO LARGE. REFER TO WARNING PLATE ON CARGO HOOK. (g) Check that only one primary ring is captured in the load beam and only one secondary ring with correct cross-section dimension is captured in the primary ring. Additional rings, slings, or shackles shall be attached to the secondary load ring (figure 4-1B). (h) Secure and clean rear view mirror, (if installed). (i) Check condition and proper length of cargo sling.

NOTE

(a) CARGO breaker ó In.

(d) Cyclic CARGO RELEASE switch ó Press and hold; pull down on cargo hook, hook should open. Release switch and cargo hook; hook should close and lock.

RELEASE

circuit

(b) BATTERY switch ó ON DC BUS 1.

c. Inflight Procedures. (1) Approach the object to be picked up with caution. A ground handler or crew member will direct the helicopter movement. (2) Maintain a constant altitude and position over the ground while the object is being placed on the cargo hook. Normally, the helicopter will be hovered into the wind. (3) After the object is secured to the cargo hook, raise the helicopter until sling is taut and lift the load

Rev. 6

4-1

BHT PUB-92-004-10

Figure 4-1. External cargo suspension system

4-2

BHT PUB-92-004-10

Figure 4-1A. Cargo release switch and decals

Rev. 6

4-2A

BHT PUB-92-004-10

Figure 4-1B. Effective loading practices

4-2B

Rev. 6

BHT PUB-92-004-10

off the ground. Takeoff will be accomplished to allow adequate clearance over all obstacles. (4) A minimum amount of control movement (to prevent oscillation of the cargo load) is desired. (5) The landing approach angle will be determined by load weight and wind condition, usually shallower than a normal approach. Do not allow the load to touch the ground until the helicopter is in a stable hover. (6) To deliver the load, lower the helicopter to relieve the tension on the sling, then use the release button or release pedal to release the load.

In order to release the cargo load safely using the cargo hook release button, the button must be depressed continuously while at lead 25 lbs. of tension is put on the cargo hook. After the load is released, the cargo hook release button must be depressed again to close the hook. NOTE The crewman does not have the capability to release the cargo load from the aft cabin. NOTE During hoist operation overtravel of the cable assembly may occur in the extended mode of operation after stopping hoist operation in MID-TRAVEL. Cable overtravel should not exceed 10 feet. If cable overtravel is observed, refer hoist to maintenance for repair. 4-2. High Performance Hoist. P r o v i s i o n s h a v e been made for the installation of an internal rescue hoist (Figure 4-2). The hoist may be installed in any one of four positions in the helicopter cabin. The hoist installation consists of a vertical column extending from floor structure to the cabin roof, a

boom with an electrically powered traction sheave, and an electrically operated winch. Two electrical control stations for the operations of the rescue hoist are provided, one for the pilot, and one for the hoist operator. A control switch is located on the cyclic control stick and provides up and down operation of the hoist as well as positioning the boom. A pendant control is provided for the hoist operator and contains a boom positioning switch and a toggle switch for hoist operation (Figure 4-3). The pilot control will override the hoist operators control. A pressure cartridge cable cutter is provided with two guarded cable cutter switches. The pilot cable cutter switch is mounted on the pedestal and the hoist operators cable cutter switch is mounted on the back of the hoist control box. The high performance hoist is an electronically speed controlled unit. Speed varies from 125 fpm at 600 pounds to 250 fpm at 300 pounds. The winch has four positive action switches. Number One is an all-stop switch that opens when three wraps of cable remain on drum. Number Two is a deceleration switch that opens when five wraps of cable remain on drum. Number Three switch has two functions, operates caution indicator light on control pendant (when caution light is on, a cable deceleration should occur) and limits cable speed when hook is 8 to 10 feet from upstow position. Number Four switch furthers limits cable speed when hook is 12 to 18 inches from the up-stow position. The first and last 20 feet of the cable are painted red. An elapsed time meter and power-on indicator are located on the control panel. A pistol grip control (Figure 4-3) is provided for the hoist operator and contains a boom in/out switch, a variable speed control, cable limit and overtemperature indicator (when hoist operating temperature limit has been exceeded the over temp. light will come on). (Secure hoist as soon as operations permit), and an intercommunication switch. The hoist has 250 feet of usable cable. Power is provided by the essential bus. Circuit protection is provided by the RESCUE HOIST POWER, RESCUE HOIST CONT, and RESCUE HOIST CABLE CUTTER circuits breakers. RESCUE HOIST CABLE CUTTER circuit breakers controls only the pilot's cable cutter switch.

Rev. 6

4-3

BHT PUB-92-004-10

Figure 4-2. Hoist Installation — Typical

4-4

Rev. 5

BHT PUB-92-004-10

Figure 4-3. Control Pendant Assembly, High Performance Hoist

Rev. 5

4-5/(4-6 blank)

BHT PUB-92-004-10

CHAPTER 5 OPERATING LIMITS AND RESTRICTIONS Section I. GENERAL 5-1.

PURPOSE.

5-3.

This chapter includes all important operating limits and restrictions that shall be observed during ground and flight operations.

5-2.

GENERAL.

The operating limitations set forth in this chapter are the direct results of design analysis, tests, and operating experiences. Compliance with these limits will allow the pilot to safely perform the assigned missions and to derive maximum utility from the helicopter. Limits concerning maneuvers, weight, and center of gravity limitations are also covered in this chapter.

EXCEEDING OPERATIONAL LIMITS.

Anytime an operational limit is exceeded an appropriate 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 that would aid maintenance personnel in the inspection that is required.

5-4.

MINIMUM CREW REQUIREMENTS.

a. The minimum crew requirement consists of a pilot, whose station is in the right seat. b. Additional crewmembers as required will be added at the discretion of the commander in accordance with applicable regulations.

Section II. SYSTEM LIMITS 5-5.

INSTRUMENT MARKINGS. (FIGURE 5-1)

a. Instrument marking color codes. Operating limitations and ranges are illustrated by the colored markings which appear on the dial faces of engine, flight, and utility system instruments. RED markings on the dial faces of these instruments indicate the limit above or below which continued operation is likely to cause damage or shorten life. The GREEN markings on the dial faces of these instruments indicate the safe or normal range of operation. The YELLOW markings on instruments indicate the range when special attention should be given to the operation covered by the instrument. Operation is permissible in the yellow range, but should be avoided. b. Instrument glass alignment marks. Limitations markings consist of strips of semitransparent color tape which adhere to the glass outside of an indicator dial. Each tape strip aligns to increment marks on the dial. Each tape strip aligns to increment marks on the dial face so correct operating limits are portrayed. The pilot should occa-

sionally verify alignment of the glass to the dial face. For this purpose, all instruments that have range marking have short, vertical white alignment marks extending from the dial glass onto the fixed base of the indicator. these slippage marks appear as a single vertical line when limitation markings on the glass properly align with reading increments on the dial face. However, the slippage marks appear as separate radial lines when a dial glass has rotated.

5-6.

ROTOR LIMITATIONS. (FIGURE 5-1)

a. The normal operating range of the main rotor is 294 to 324 rpm and the range is marked on the dual tachometer as a green arc on the face of the instrument. b. Normally, autorotation rpm will be set at approximately 295 to 324 rotor rpm at seal level, 70 KIAS, at light gross weight (2 persons, full fuel). Autorotation main rotor speed shall not exceed 339 rpm. c. Main rotor speeds in excess of 339 rpm shall be entered in DA Form 2408-13.

Rev. 5

5-1

BHT PUB-92-004-10

Figure 5-1. Instrument markings (Sheet 1 of 2)

5-2

Rev. 6

BHT PUB-92-004-10

Figure 5-1. Instrument markings (Sheet 2 of 2)

Rev. 7

5-3

BHT PUB-92-004-10

d. Rotor Operating Limits decal, located at the lower left of the instrument panel, specifies limiting conditions at specific altitudes and gross weights. e. It is possible to encounter blade stall within the operatin range under high gross weight, high altitude, or high airspeed, or during acceleration or low rpm. Blade stall and the remedy are more thoroughly discussed in Chapter 8. f. At heavy gross weights, high density altitudes, or during maneuvering, the rotor rpm will tend to overspeed and shall be monitored and

controlled by the pilot, using collective pitch to keep the rotor within limits. g. Wind Limitations. Helicopter can be started in a maximum wind velocity of 30 knots and a maximum gust spread of 15 knots. NOTE Gust spreads are not normally reported. To obtain spread, compare minimum and maximum wind velocity. h. Rotor Brake Operating Limits. The rotor brake shall not be applied above 130 rotor RPM.

Section III. POWER LIMITS 5-7.

ENGINE LIMITATIONS.

d. Maximum oil consumption is 0.3 gal (2.4 pints) per hour.

5-7A. STARTER LIMITATIONS. An entry in DA Forms 2408-13 is required if the following limits are exceeded. a. Maximum N2 rpm allowable is 6840 at or below 91% N1.

Three energized periods allowed per hour. Limit starter energizing time to: 35 seconds - ON. 3 minutes - OFF. 35 seconds - ON.

b. Maximum N2 rpm allowable is 6640 above 91% N1.

30 minutes - OFF. 35 seconds - ON.

5-8. If transient N2 limits are reached, main rotor RPM limits have been exceeded. A main rotor overspeed inspection must be performed in accordance with BHT PUB92-004-23. c. Maximum transient N2 rpm allowable is 6910 for a maximum of three seconds.

5-4

Rev. 7

ENGINE RATING.

The helicopter is equipped with a T53-L-703 series engine which at 6600 rpm has the following rates: Take-Off Power (TOP) — 1800 shp and Maximum Continuous Power (MCP) — 1500 shp. The helicopter is torque limited by the transmission to TOP (5 minutes) — 1290 shp and MCP — 1134 shp which corresponds to 100% and 88% on the torque gauge when operating at 6600 engine/324 rotor rpm.

BHT PUB-92-004-10

Section IV. LOADING LIMITS 5-9.

CENTER OF GRAVITY LIMITATIONS.

a. Flight. Center of gravity (CG) limits for loading purposes are located between fuselage stations 130 and 144. (Refer to Chapter 6 for CG limits chart.)

CG limits may be exceeded under minimum payload and crew conditions when either the armor seats are removed or if the battery is in the aft location. Check to ensure helicopter is within limits prior to flight.

Do not carry external loads with a CG aft of station 142 prior to lifting external load.

When flying at an aft CG (station 140 to 144) terminate an approach at a minimum of 5-foot hover prior to landing to prevent striking the tail on the ground. Practice touchdown autorotations shall not be attempted with the CG aft of station 140 because termination at 5 feet is not possible.

5-10. WEIGHT LIMITATIONS. NOTE Actual weight change shall be determined after cargo hook kit is installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits. a. Maximum Gross Weight. The maximum gross weight for the helicopter is 10,500 pounds for internal loading and 11,200 pounds for external loading. The maximum gross weights for varying conditions of temperature, altitude, wind velocity, and skid height is shown in Chapter 7. b. Maximum Gross Weight for Towing. Towing the helicopter, with ground handling wheels installed, on rough surfaces at gross weights in excess of 9,500 pounds may cause permanent set in the aft cross tube. c. Cargo Limitations.

Hook

P/N

205-070-900

Weight

Move battery to the forward location when operating with bladder-type ferry fuel to remain within CG limits.

Due to the danger of high voltage created by static electricity in e x t e rn a l c a r g o h o o ku p s , a c r e w member must inform al l ground handling personnel of high voltage hazards when making external cargo hookups. Caution shall be exercised when carrying external cargo because the helicopter handling characteristics may be affected by the size, weight, and shape of the cargo load. Maximum allowable weight of external cargo is 4000 pounds.

b. Crew/Passenger Shift. For information pertaining to CG shift due to crew movement, refer to figure 5-2.

d. Weight Distribution Limitations. For information pertaining to weight distribution and helicopter performance. (Chapters 6 and 7.)

c. Lateral Center of Gravity Limitations. The lateral CG limits are plus or minus 7.5 inches. For additional information on computation of lateral CG, refer to Chapter 6.

With armored seats installed, and with a left lateral CG the pilots arm and right cyclic movement will be restricted.

e. Minimum Weight. Refer to Chapter 6. f. Floor Loading Limits. High density cargo distributed over the deck area to maintain 100 pounds per square foot will provide a safety load factor of 3.0 based on load limit. (Refer to Chapter 6.) g. Gross Weight for Safe Pedal Margin Chart. Refer to figure 5-3. h. Fuel Versus Alternate Load Capacity Envelope. Refer to Chapter 7.

Rev. 6

5-5

BHT PUB-92-004-10

Figure 5-2. Crew movement

5-6

BHT PUB-92-004-10

c. Cargo Limitations.

Hook

P/N

212-706-103

Weight

d. Weight Distribution Limitations. For information pertaining to weight distribution and helicopter performance. (Chapters 6 and 7.) e. Minimum Weight. Refer to Chapter 6.

Due to the danger of high voltage created by static electricity in external cargo hookups, a crew member must inform all ground handling personnel of high voltage hazards when making external cargo hookups. Caution shall be exercised when carrying external cargo because the helicopter handling characteristics may be affected by the size, weight, and shape of the cargo load. Maximum allowable weight of external cargo is 5000 pounds.

f. Floor Loading Limits. High density cargo distributed over the deck area to maintain 100 pounds per square foot will provide a safety load factor of 3.0 based on load limit. (Refer to Chapter 6.) g. Gross Weight for Safe Pedal Margin Chart. Refer to figure 5-3. h. Fuel Versus Alternate Load Capacity Envelope. Refer to Chapter 7.

Rev. 6

5-6A/(5-6B blank)

BHT PUB-92-004-10

Figure 5-3. Gross weight for safe pedal margin chart

5-7

BHT PUB-92-004-10

Section V. AIRSPEED LIMITS modifications have been made to the cargo doors and airframe (figure 5-4).

5-11. AIRSPEED LIMITATIONS.

a. Refer to Chapter 7 for forward airspeed limits. If a main cargo door comes open from closed position while in flight reduce airspeed below 50 knots and secure door. If not possible to secure in closed position, the door must be fully opened and secured by the open door latch (figure 54). When securing the cargo door in flight, the crew members shall be fastened to the helicopter by seat belts or other safety devices.

The helicopter can be flown at 120 KIAS with the cargo doors locked in the full open position only if the appropriate

b. Operation In Takeoff Power Range. Vne is 80 KIAS for all gross weights when above maximum continuous torque (88%). c. Sideward flight limits are 35 knots. d. Rearward flight limit is 30 knots. e. Climb. During climbs at low altitude, safe autorotative speed shall be maintained in the event an engine failure occurs, sufficient airspeed is available to accomplish a safe autorotative landing. Refer to Paragraph 5-14 and applicable chart for details concerning climb limitations.

Figure 5-4. Open door latch

5-8

Rev. 7

BHT PUB-92-004-10

Section VI. MANEUVERING LIMITS 5-12. PROHIBITED MANEUVERS.

Abrupt changes of flight controls causing a negative "G" load are prohibited.

a. Protracted rearward flight and downwind hovering are prohibited. b. The speed for any and all maneuvers shall not exceed the level flight velocities as stated on the Air Speed Operating Limits Chart (figure 5-5).

Section VII. ENVIRONMENTAL RESTRICTIONS 5-13. ENVIRONMENTAL RESTRICTIONS. This helicopter is qualified for flight under instrument meteorological conditions, except as noted in (Chapter 8) Normal Procedures.

Section VIII. HEIGHT VELOCITY 5-14. HEIGHT VELOCITY. Refer to the Minimum Height For Safe Landing After Engine Failure Chart.

Rev. 1

5-9

BHT PUB-92-004-10

Figure 5-5. Airspeed operating limits chart

5-10

BHT PUB-92-004-10

Figure 5-6. Minimum height for safe landing after engine failure chart

5-11/(5-12 blank)

BHT PUB-92-004-10

CHAPTER 6 WEIGHT/BALANCE AND LOADING

Section I. GENERAL 6-1.

GENERAL.

This chapter contains sufficient instructions and data so that an aviator knowing the basic weight and moment of the helicopter can compute any combination of weight and balance.

6-2.

CLASSIFICATION OF HELICOPTER.

For the purpose of clarity, the Bell UH-1H-II helicopter is in Class I. Additional directives

governing weight and balance of Class I helicopter forms and records, are contained in AR95-16 and TM 1500-342-23.

6-3.

HELICOPTER STATION DIAGRAM.

Figure 6-1 shows the helicopter reference datum lines, fuselage stations, buttlines, waterlines and jack pad locations. The primary purpose of the figure is to aid personnel in the computation of helicopter weight / balance and loading.

Section II. WEIGHT AND BALANCE 6-4.

LOADING CHARTS.

a. Information. The loading data contained in this chapter is intended to provide information necessary to work a loading problem for the helicopter to which this manual is applicable. b. Use. From the figures contained in this chapter, weight and moment are obtained for all variable load items and are added to the current basic weight and moment (DD Form 365-3) to obtain the gross weight moment. (1) The gross weight and moment are checked on figure 6-2 to determine the approximate center of gravity (CG). (2) The effect on CG by the expenditures in flight of such items as fuel may be checked by subtracting the weights and moments of such items from the takeoff weight and moment and checking the new weight and moment on the Loading Limits Chart. (3) If the weight and moment lines do not intersect, the CG is not within the flight limits. NOTE This check should be made to determine whether or not the CG will remain within limits during the entire flight.

6-5. DD FORM 365-1 — BASIC WEIGHT CHECKLIST. The form is initially prepared by the manufacturer before the helicopter is delivered. The form is a tabulation of equipment that is, or may be, installed and for which provision for fixed stowage has been made in a definite location. The form gives the weight, arm, and moment /100 of individual items for use in correcting the basic weight and moment on DD Form 365-3 as changes are made in this equipment.

6-6. DD FORM 365-3 — BASIC WEIGHT AND BALANCE RECORD. The form is initially prepared by the manufacturer at time of delivery of the helicopter. The form is a continuous history of the basic weight and moment resulting from structural and equipment changes. At all times the last entry is considered current weight and balance status of the basic helicopter. Figure 6-2 shows a sample DD Form 365-3.

6-7. DD FORM 365-4 — WEIGHT AND BALANCE CLEARANCE FROM F TRANSPORT AND HELICOPTER. a. General. The form is a summary of actual disposition of the load in the helicopter. It records the balance status of the helicopter, step-by-step. It serves as a worksheet on which to record weight and

Rev. 1

6-1

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Figure 6-1. Helicopter station diagram

6-2

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Figure 6-2. DDS Form 365-3

6-3

BHT PUB-92-004-10 balance calculations, and any corrections that must be made to ensure that the helicopter will be within weight and CG limits. The form is required only when the loading is such as to seriously affect the flying characteristics and safety of the helicopter and in all cases where alternate loading is employed. b. Form Preparations. Specific instructions for filling out the form is given in the following paragraphs. Figure 6-3 (sheet 1) shows the results of the instructions.

(10)Reference 9 ó Enter the sum of the weights of reference 1 through 7 inclusive to obtain "operating weight" (11)Reference 10 ó Enter the number of gallons and weight of "Take-Off Fuel". (12)Reference 11 ó Not applicable. (13)Reference 12 ó Enter the sum of the weights for reference 8 through 10 inclusive to obtain "Total Helicopter Weight". (14)Determine the "Allowable Load" based on takeoff and landing by use of the "Limitations" table in the lower left-hand corner of the form as follows:

NOTE Reference 1, 2, etc., are references to items 1, 2, etc., on DD Form 365-4.

(a) Enter the "Allowable Gross Weight" for takeoff and landing.

(1) Insert the necessary identifying information at the top of the form. In the blank spaces of the "Limitation" table, enter the gross weight and CG restrictions obtained from Chapter 5.

(b) Enter the "Total Helicopter Weight" (from reference 12). Estimate the fuel to be aboard at time of landing. Enter the sum "Operating Weight" (from reference 9) and "Estimated Landing Fuel Weight".

(2) Reference 1 ó Enter the helicopter basic weight and moment from the last entry on Chart C ó Basic Weight and Balance Record.

(c) Subtract above weights from the respective "Allowable Gross Weight" to obtain the respective "Allowable Loads". NOTE

NOTE Enter moment/100 values throughout the form. Obtain these values form this chapter. (3) Reference 2 ó Oil included in basic weight. (4) Reference 3 ó enter the number and weight of crew. Use actual crew weight, if available. (5) Reference 4 ó Enter the weight of the crew baggage, if applicable. (6) Reference 5 ó Not applicable. (7) Reference 6 ó Enter the emergency equipment, if applicable.

weight of

(8) Reference 7 ó Enter the weight of any extra equipment, if applicable. (9) Reference 8 blank.

6-4

The smallest of these allowable loads is the "Allowable Load" and represents the maximum amount of weight which may be distributed throughout the helicopter in the various compartments without exceeding the maximum gross weights of the helicopter. (15)Reference 13 ó Using the helicopter diagram enter the number and weight of passengers and the weight of cargo (baggage). Use actual Passenger Weight, if available. Enter the total for each compartment in the weight column. NOTE NOTE The sum of the compartment totals must not exceed the "Allowable Load" determined in the "Limitations Table". (16)Reference 16 ó Enter the sum of reference 12 and the compartment totals from reference 13 opposite "Takeoff Condition" (uncorrected). At this point, if not already done, calculate the moment/100 for reference 1 through 16 inclusive.

BHT PUB-92-004-10

Figure 6-3. DD Form 365-4 (Sheet 1 of 2)

6-5

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Figure 6-3. DD Form 365-4 (Sheet 2 of 2)

6-6

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(17)Check the weight figure (reference 16) against the "Gross Weight Takeoff" in the "Limitations" table. Check the weight and moment/ 100 figure opposite reference 16 on the Loading Limit Chart to ascertain that the indicated CG is within allowable limits. (18)Reference 18 ó If changes in amount or distribution of load are required, indicate necessary adjustments by proper entries in the "Corrections" table left-hand side of the form, as follows: (a) Enter a brief description of the adjustment made in the column marked "Item". (b) Add all the weight and moment decreases and insert the totals in the space opposite "Total Gross Weight". (c) Add all the weight and moment increases and insert the total in the space opposite "Total Weight Added". (d) Subtract the smaller from the larger of the two totals and enter the difference (with applicable plus or minus sign) opposite "Net difference". (e) Transfer these "Net Difference" figures to the space opposite reference 18. (19)Reference 19 ó Enter the sum of the difference between reference 16 and reference 18. Recheck to see that these figures do not exceed allowable limits. (20)Reference 20 ó Determine the takeoff CG position by referring to the Loading Limit chart. Enter this figure in the space provided opposite "Takeoff CG". (21)Reference 21 ó Enter the operating weight from reference 9 which is the zero FUEL WEIGHT AND MOMENT. (22)Reference 22 ó Enter the weight of "Air Supply Load" to be dropped before landing with moment /100 if applicable.

ascertain that the CG is within the allowable limits. If not, make necessary changes to landing fuel and recalculate (21) through (25).

6-8. DD FORM 365-4 WEIGHT AND BALANCE CLEARANCE FORM F TACTICAL HELICOPTERS. a. General. The form is a summary of actual disposition of the load in the helicopter. It records the balance status of the helicopter, step-by-step. It serves as a worksheet on which to record weight and balance calculations, and any corrections that must be made to ensure that the helicopter will be within weight and CG limits. The form is required only when the loading is such as to seriously affect the flying characteristics and safety of the helicopter and in all cases where alternate loading is employed. b. Form Preparation. Specific instructions for filling out the form is given in the following paragraphs. Figure 6-3 (sheet 2) shows the results of the instructions. NOTE Reference 1, 2, etc., are references to items 1, 2, etc., on DD Form 365-4. (1) Insert the necessary identifying information at the top of the form. In the blank spaces of the "Limitations" table enter the gross weight and CG restrictions obtained from Chapter 5. NOTE Enter moment /100 values throughout the form. Obtain these values from this chapter. (2) Reference 1 ó Enter the helicopter basic weight and moment /100. Obtain these figures from the last entry on Chart C ó Basic Weight and Balance Record.

Enter estimated landing

(3) Reference 2 ó OIL INCLUDED IN BASIC WEIGHT.

(24)Reference 24 ó Total items 21, 22 and 23. This is the estimated landing condition.

(4) Reference 3 ó Enter the number and weight of the crew at their take-off stations. Use actual crew weights if available. Also, enter the weight of baggage, cargo, and miscellaneous items. Enter the total of each compartment in the "Weight" column.

fuel.

(23)Reference 23 ó

(25)Reference 25 ó Check the weight and moment /100 figures on the Loading Limits chart to

Rev. 1

6-7

BHT PUB-92-004-10 (5) Reference 4 ó Enter the sum of the weights for references 1 through 3 to obtain "Operating Weight". (6) Reference 5 ó Enter the number of rounds, caliber, and weight of all ammunition of applicable. (7) Reference 6 ó Enter the size, distribution (forward, aft, external, etc.), and weight of all bombs, torpedoes, rockets, etc., if applicable. (8) Reference 7 ó Enter the number of gallons and weight of fuel. If auxiliary fuel tanks are to be used, these items and their weight should also be entered as part of reference 7. (9) Reference 8 ó Not applicable. (10)Reference 9 ó Not applicable. (11)Reference 10 ó Enter the sum of the weights for references 4 through 9 opposite "TakeOff Condition" (Uncorrected). At this point, if not already done, calculate and enter the moment/100 for references 1 through 10. (12)Check the weight figure opposite reference 10 against the "Gross Weight Take-Off" in the "Limitations" table. check the moment/100 figure opposite reference 10 to verify that the indicated CG is within allowable limits in the "Limitations" table. (13)Reference 11 ó If changes in weight or distribution of load are required, indicate necessary adjustments by prior entries in the "Corrections" table in lower left-hand corner of the form as follows: (a) Enter a brief description adjustment in the column marked "Item".

of

the

(b) Add all the weights and add all the moment decreases. Insert the totals in the space opposite "Total Weight Removed". (c) Add all the weights and add the moment increases. Insert the totals in the space opposite "Net Difference". (d) Subtract the smaller from the larger of the two totals and enter the difference (with applicable plus or minus sign) opposite "Net Difference".

6-8

(e) Transfer these net difference figures to the spaces opposite reference 11. (14)Reference 12 ó Enter the sum of, or the difference between, reference 10 and 11. Recheck to verify that these figures do not exceed allowable limits. (15)Reference 13 ó By referring to the CG table, determine the take-off CG position. Enter this figure in the space provided opposite "Take-Off CG". (16)Reference 14 ó Estimate the weights of ammunition (not including weight of cases and links if retained), fuel, paratroopers (use actual weight of troops with all equipment, if available), external cargo, and any other items which may be expended before landing. Enter these figures together with their moment/100 in the space provided.

NOTE Do not consider reserve fuel as expended when determining "Estimate Landing Condition". (17)Reference 15 ó Enter the difference in weights and moment/100 between reference 12 and the total of reference 14. (18)Reference 16 ó By again referring to the CG table, determine the estimated landing CG position. Enter the figure opposite "Estimated Landing CG".

NOTE Check the landing CG figure with permissible CG figures in limitation block. The landing CG shall be within the range of the figures shown. (19)The necessary signatures shall appear at the bottom of the form.

NOTE For charts and forms refer to Weight and Balance Control Data, Military Specification MIL-W-25140B.

BHT PUB-92-004-10

Section III. PERSONNEL 6-9.

single-passenger folding seats, facing aft with backs, are located aft of the pilot and copilot seats.

TROOP SEATS.

Refer to figure dimensions.

6-4

for

cargo

compartment NOTE

6-10. TROOP SEATS. a. Description. The troop seats are of tubular construction with reinforced canvas webbing for support areas. The seats are attached to the floor and transmission support structure. Seats can be installed for rescue missions, then folded and stowed flat; or they can be folded for cargo missions as required. b. Arrangement. eleven passengers can be seated in the aft area for the forward fuselage section. Either of the two following arrangements may be used for passenger seating (figure 6-5 and 6-6). (1) Three seats facing forward, and accommodating five passengers, may be placed across the cabin immediately forward of the transmission support structure. A one-passenger seat, without back rest, is located between two-man seats which have backs. Two two-man seats, without backs, are located aft of the five-passenger seats parallel to the helicopter center line. Passengers in these seats face outboard. Two single passenger folding seats, facing aft with backs, are located just aft of the crew seats. (2) Four two-man seats, facing outboard, may be placed, two on each side of the helicopter center line, approximately in line with the side faces of the transmission support structure. The two forward seats are equipped with backs. A one-passenger seat, without back rest, is located immediately forward of the transmission support structure on the helicopter center line and faces forward. Two

The two single passenger seats located aft of the pilot and copilot seats may be removed and one facing forward installed in the center position directly aft of console. c. Troop Seat Belts. Individual lap-type seat belts are provided for all troop seats. These same belts, with web extensions, are provided for litter patients when helicopter is used for rescue missions.

6-11. PERSONNEL LOADING AND UNLOADING. When helicopter is operated at critical gross weights, the exact weight of each individual occupant plus equipment should be used. If weighing facilities are not available, or if the tactical situation dictates otherwise, loads shall be computed as follows: a. Combat individual.

equipped

soldiers:

240

lb

per

b. Combat equipped paratroopers: 260 lb per individual. c. Crew and passengers with no equipment: compute weight according to each individuals estimate.

6-12. PERSONNEL MOMENTS. Refer to figure 6-7.

Section IV. MISSION EQUIPMENT (Not Applicable)

6-9

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Figure 6-4. Cargo compartment

6-10

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Figure 6-5. Troop seat placement

6-11

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Figure 6-6. Alternate troop seat placement

6-12

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Figure 6-7. Personnel moment (Sheet 1 of 2)

6-13

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Figure 6-7. Personnel moment (Sheet 2 of 2)

6-14

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Section V. CARGO LOADING 6-13. CARGO LOADING. a. The large cargo doors, open loading area and low floor level preclude the need for special loading aids. Through loading may be accomplished by securing cargo doors in the fully open position. Cargo tiedown fittings (figure 6-8 and 6-9) are located on the cabin floor for securing cargo to prevent cargo shift during flight. b. Preparation of General Cargo. (1) The loading crew shall assemble the cargo and baggage to be transported. At time of assembly and prior to loading, the loading crew shall compile data covering weight, dimensions, center of gravity location and contact areas for each item. (2) Heavier packages to be loaded shall be loaded first and placed in the aft section against the bulkhead for CG range purposes. Helicopter floor loading in this area shall not exceed 100 pounds per square foot maximum package size and gross weight limits. The baggage compartment loading shall not exceed 400 pounds maximum. (3) Calculation of the allowable load and loading distribution shall be accomplished by determining the final CG location and remain within the allowable limits for safe operating conditions. (4) A loading chart is located on the right hand hinged door post.

6-14. CARGO CENTER OF GRAVITY PLANNING. a. The items to be transported should be assembled for loading after the weight and dimensions have been recorded. (1) Loading time will be gained if the packages are positioned as they are to be located in the helicopter. (2) To assist in determining the locations of the various items, the individual weights and total weight must be known. (3) When these factors are known the Internal, External Cargo and Baggage Loading (figure 6-10, 611 and 6-12) can be used as a guide to determine the helicopter station at which the package CG shall be located. Baggage compartment is accessible from right side of tailboom and contains approximately 28 cubic feet of space. Baggage compartment has a load limit of 400 pounds (181.4 kilograms) not to exceed 100 pounds per square foot (0.048 kilograms per square centimeter). These are structural limitations only and do not infer that CG will remain within

approved limits. When weight is loaded into baggage compartment, indiscriminant crew, passenger, and fuel loading can no longer be assumed, and pilot must compute GW CG to ensure loading is within approved limits. (4) The information presented on the loading chart will not be affected by fuel quantity, as full to empty fuel load has been considered during data computation. (5) Final analysis of helicopter CG location for loading shall be computed from the data presented in this chapter. b. Computation of Cargo Center of Gravity. NOTE Actual weight change shall be determined after cargo hook kit is installed and ballast readujsted, if necessary, to return empty weight CG to within allowable limits. (1) The loading data in this chapter will provide information to work a loading problem. From the loading tables, weight and moment/100 are obtained for all variable load items and are added mathematically to the current basic weight and moment/100 obtained from Chart C to arrive at the gross weight and moment. (2) The CG of the loaded helicopter is represented by a moment figure in the center of gravity table. If the helicopter is loaded within the forward and aft CG limits, the figure will fall numerically between the limiting moments. (3) The effect on the CG of the usable inflight items of fuel and oil may be checked by subtracting the weights and moments of such items from the takeoff gross weights and moments of such items from the takeoff gross weight and moment and checking the new moment with the CG table. (4) This check will be made to determine whether or not the CG will remain within limits during the entire flight. c. Loading Procedures. The helicopter requires no special loading preparation. (1) The loading procedure consists of placing the heaviest items to be loaded as far aft as possible. Such placement locates the cargo nearer the helicopter CG and allows maximum cargo load to be transported, as well as maintaining the helicopter within the safe operating CG limits for flight. (2) The mission to be performed should be known to determine the cargo, troop transport, or litter patients to be carried on the return trip. Rev. 6

6-15

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Figure 6-8. Cargo tiedown fitting data

6-16

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Figure 6-9. Cargo loading typical

Rev. 1

6-17

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Figure 6-10. Internal loading

6-18

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Figure 6-11. External loading

Rev. 6

6-19

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Figure 6-12. Baggage compartment loading

6-20

Rev. 2

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(3) If troops or litter patients are to be carried, troop seats and litter racks shall be loaded aboard and stowed. d. Securing Loads. Equipment for securing cargo consists of a three-piece cargo tiedown net, which attaches to tiedown rings. e. Cargo Loading ó Internal.

will provide a safety load factor of 3.0 based on limit loads. (4) The safety load factor will vary as the floor loading varies. (Load factor = 300 pounds per square foot floor loading.) Flush-mounted tiedown fittings are provided on the beam and aft cabin bulkhead.

6-15. LITTER RACKS.

(1) Internal cargo is carried within the cabin, and bulk items can be accommodated. (2) The cargo area contains approximately 220 cubic feet of obstruction-free cargo load space. (3) High density cargo distributed over the deck area to maintain 100 pounds per square foot

The litter rack installation (figure 6-13), accommodates three stretchers parallel to the font edge of the transmission support structure. They can be installed for transporting litter patients or removed for carrying cargo or personnel. Two medical attendant seats are attached to the cabin floor forward of the litters, facing aft.

Section VI. FUEL 6-16. FUEL. Refer to figure 6-14.

Section VII. ALLOWABLE LOADING 6-17. ALLOWABLE LOADING. For longitudinal gross weight limits refer to figure 6-14A.

The takeoff gross weight at sea level on a standard day 15∞C (59∞F) 29.92 inches Hg is 10,500 lbs. Refer to figure 6-15.

Rev. 6

6-21

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Figure 6-13. Litter installation — typical

6-22

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Figure 6-14. Fuel data (Sheet 1 of 4)

6-23

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Figure 6-14. Fuel data (Sheet 2 of 4)

6-24

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Figure 6-14. Fuel data (Sheet 3 of 4)

Rev. 5

6-25

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Figure 6-14. Fuel data (Sheet 4 of 4)

6-26

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Figure 6-14A. Longitudinal gross weight limits

Rev. 6

6-26A/(6-26B blank)

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Figure 6-15. Allowable loading (Sheet 1 of 3)

Rev. 1

6-27

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Figure 6-15. Allowable loading (Sheet 2 of 3)

6-28

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Figure 6-15. Allowable loading (Sheet 3 of 3)

Rev. 1

6-29/(6-30 blank)

BHT PUB-92-004-10

CHAPTER 7 PERFORMANCE DATA Section I. INTRODUCTION 7-1.

PURPOSE.

The purpose of this chapter is to provide the best available performance data for the UH-1H-II helicopter. Regular use of this information will enable you to receive maximum safe utilization from the helicopter. Although maximum performance is not always required, regular use of this chapter is recommended for the following reasons:

d. Experience will be gained in accurately estimating the effects of variables for which data are not presented. NOTE The information provided in this chapter is primarily intended for mission planning and is most useful when planning operations in unfamiliar areas or at extreme conditions. The data may also be used inflight to establish unit or area standing operating procedures and to inform ground commanders of performance risk tradeoffs.

a. Knowledge of your performance margin will allow you to make better decisions when unexpected conditions or alternate missions are encountered. b. Situations requiring maximum performance will be more readily recognized.

7-2.

c. Familiarity with the data will allow performance to be computed more easily and quickly.

The following index contains a list of the sections and their titles, the figure numbers, subjects and page numbers of each performance data chart contained in this chapter.

CHAPTER 7 INDEX.

INDEX Section

Page No.

Subject

I

Introduction......................................................................................................................

7-1

II

Torque Available...............................................................................................................

7-7

Figure 7-1. Maximum Torque (5 Minute Operation) Chart 324 Rotor/6600 Engine RPM..................................................................................

7-8

III

Hover

7-9

Figure 7-1A. Hover Torque Required ......................................................................... Figure 7-2. Hover ceiling in ground (IGE) (5 Minute Operation) - Heater On ........ Figure 7-2A. Hover ceiling in ground (IGE) (5 Minute Operation) - Heater Off...... Figure 7-3. Hover ceiling in ground (IGE) (Continuous Operation) - Heater On .... Figure 7-3A. Hover ceiling in ground (IGE) (Continuous Operation) - Heater Off.. Figure 7-3B. Hover ceiling out of ground (OGE) (5 Minute Operation) - Heater On Figure 7-3C. Hover ceiling out of ground (OGE) (5 Minute Operation) - Heater Off Figure 7-3D. Hover ceiling out of ground (OGE) (Continuous Operation) - Heater On Figure 7-3E. Hover ceiling out of ground (OGE) (Continuous Operation) - Heater Off Figure 7-3F. Critical relative wind azimuths for hover flight ................................... IV

Control Margin ................................................................................................................. Figure 7-4. Control Margin Chart Sheet 1 of 2 .......................................................... Figure 7-4. Control Margin Chart Sheet 2 of 2 .......................................................... Rev. 7

7-10 7-11 7-12 7-13 7-14 7-14A 7-14B 7-14C 7-14D 7-14E 7-15 7-16 7-17 7-1

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INDEX (Cont) Section V

VI

7-2

Subject Takeoff .............................................................................................................................. Figure 7-5. Takeoff Chart,, Sheet 1 of 3, Level Acceleration (3 Foot Skid Height)............................................. Sheet 2 of 3, Climb and Acceleration (3 Foot Skit Height) ..................................... Sheet 3 of 3, Level Acceleration (15 Foot Skid Height)........................................... Cruise................................................................................................................................ Figure 7-6. Cruise Charts. Sheet 1 of 48, FAT = –30°C, Pressure Altitude Seal Level ..................................... Sheet 2 of 48, FAT = –30°C, Pressure Altitude 2000 FT......................................... Sheet 3 of 48, FAT = –30°C, Pressure Altitude 4000 FT......................................... Sheet 4 of 48, FAT = –30°C, Pressure Altitude 6000 FT......................................... Sheet 5 of 48, FAT = –30°C, Pressure Altitude 8000 FT......................................... Sheet 6 of 48, FAT = –30°C, Pressure Altitude 10000 FT....................................... Sheet 7 of 48, FAT = –30°C, Pressure Altitude 12000 FT....................................... Sheet 8 of 48, FAT = –30°C, Pressure Altitude 14000 FT....................................... Sheet 9 of 48, FAT = –15°C, Pressure Altitude Seal Level ..................................... Sheet 10 of 48, FAT = –15°C, Pressure Altitude 2000 FT....................................... Sheet 11 of 48, FAT = –15°C, Pressure Altitude 4000 FT....................................... Sheet 12 of 48, FAT = –15°C, Pressure Altitude 6000 FT....................................... Sheet 13 of 48, FAT = –15°C, Pressure Altitude 8000 FT....................................... Sheet 14 of 48, FAT = –15°C, Pressure Altitude 10000 FT..................................... Sheet 15 of 48, FAT = –15°C, Pressure Altitude 12000 FT..................................... Sheet 16 of 48, FAT = –15°C, Pressure Altitude 14000 FT..................................... Sheet 17 of 48, FAT = 0°C, Pressure Altitude Seal Level ....................................... Sheet 18 of 48, FAT = 0°C, Pressure Altitude 2000 FT........................................... Sheet 19 of 48, FAT = 0°C, Pressure Altitude 4000 FT........................................... Sheet 20 of 48, FAT = 0°C, Pressure Altitude 6000 FT........................................... Sheet 21 of 48, FAT = 0°C, Pressure Altitude 8000 FT........................................... Sheet 22 of 48, FAT = 0°C, Pressure Altitude 10000 FT......................................... Sheet 23 of 48, FAT = 0°C, Pressure Altitude 12000 FT......................................... Sheet 24 of 48, FAT = 0°C, Pressure Altitude 14000 FT......................................... Sheet 25 of 48, FAT = 15°C, Pressure Altitude Seal Level ..................................... Sheet 26 of 48, FAT = 15°C, Pressure Altitude 2000 FT......................................... Sheet 27 of 48, FAT = 15°C, Pressure Altitude 4000 FT......................................... Sheet 28 of 48, FAT = 15°C, Pressure Altitude 6000 FT......................................... Sheet 29 of 48, FAT = 15°C, Pressure Altitude 8000 FT......................................... Sheet 30 of 48, FAT = 15°C, Pressure Altitude 10000 FT....................................... Sheet 31 of 48, FAT = 15°C, Pressure Altitude 12000 FT....................................... Sheet 32 of 48, FAT = 15°C, Pressure Altitude 14000 FT....................................... Sheet 33 of 48, FAT = 30°C, Pressure Altitude Seal Level ..................................... Sheet 34 of 48, FAT = 30°C, Pressure Altitude 2000 FT......................................... Sheet 35 of 48, FAT = 30°C, Pressure Altitude 4000 FT......................................... Sheet 36 of 48, FAT = 30°C, Pressure Altitude 6000 FT......................................... Sheet 37 of 48, FAT = 30°C, Pressure Altitude 8000 FT......................................... Sheet 38 of 48, FAT = 30°C, Pressure Altitude 10000 FT.......................................

Page No. 7-19 7-21 7-22 7-23 7-25 7-27 7-29 7-31 7-32 7-33 7-34 7-35 7-36 7-37 7-38 7-39 7-40 7-41 7-42 7-43 7-44 7-45 7-46 7-47 7-48 7-49 7-50 7-51 7-52 7-53 7-54 7-55 7-56 7-57 7-58 7-59 7-60 7-61 7-62 7-63 7-64 7-65 7-66

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INDEX (Cont) Section

VII

VIII

IX

X

XI XII XIII XIV XV

Page No.

Subject Sheet 39 of 48, FAT = 30∞C, Pressure Altitude 12000 FT....................................... Sheet 40 of 48, FAT = 30∞C, Pressure Altitude 14000 FT....................................... Sheet 41 of 48, FAT = 45∞C, Pressure Altitude Sea Level ...................................... Sheet 42 of 48, FAT = 45∞C, Pressure Altitude 2000 FT......................................... Sheet 43 of 48, FAT = 45∞C, Pressure Altitude 4000 FT......................................... Sheet 44 of 48, FAT = 45∞C, Pressure Altitude 6000 FT......................................... Sheet 45 of 48, FAT = 45∞C, Pressure Altitude 8000 FT......................................... Sheet 46 of 48, FAT = 55∞C, Pressure Altitude Sea Level ...................................... Sheet 47 of 48, FAT = 55∞C, Pressure Altitude 2000 FT......................................... Sheet 48 of 48, FAT = 55∞C, Pressure Altitude 4000 FT.........................................

7-67 7-68 7-69 7-70 7-71 7-72 7-73 7-74 7-75 7-76

Drag ..................................................................................................................................

7-77

Figure 7-7. Drag Chart................................................................................................ Sheet 1 of 2, Drag (Authorized Configurations) Chart ........................................... Sheet 2 of 2, Drag Chart ...........................................................................................

7-79 7-81

Climb Performance/Climb-Descent .................................................................................

7-83

Figure 7-8. Climb Performance Chart ........................................................................ Figure 7-9. Climb-Descent Chart................................................................................

7-85 7-86

Fuel Flow ..........................................................................................................................

7-87

Figure 7-10. Idle Fuel Flow......................................................................................... Figure 7-11. Fuel Flow Vs Torque, 0∞C ......................................................................

7-88 7-89

Nautical Miles and Time and Range...............................................................................

7-90

Figure 7-12. Nautical Miles Per Pound of Fuel.......................................................... Figure 7-13. Time and Range Vs Fuel........................................................................

7-91 7-93

Autorotational Glide Characteristics ..............................................................................

7-94

Figure 7-14. Autorotation Glide Characteristics .......................................................

7-95

Airspeed Limits ................................................................................................................

7-96

Figure 7-15. Airspeed Operating Limits ....................................................................

7-97

Airspeed Correction .........................................................................................................

7-98

Figure 7-16. Airspeed Correction Chart .....................................................................

7-99

Temperature .....................................................................................................................

7-100

Figure 7-17. Temperature Conversion Chart.............................................................

7-101

Performance Planning .....................................................................................................

7-102

Figure 7-18. Performance Planning Card ..................................................................

7-104

Rev. 1

7-3

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7-3.

GENERAL.

The data presented covers the maximum range of conditions and performance that can reasonably be expected. In each area of performance, the effects of altitude, temperature, gross weight, and other parameters relating to that phase of flight are presented. In addition to the presented data, your judgement and experience will be necessary to accurately obtain performance under a given set of circumstances. The conditions for the data are listed under the title of each chart. The effects of different conditions are discussed in the test accompanying each phase of performance. Where practical, data are presented at conservative conditions. however NO GENERAL CONSERVATISM HAS BEEN APPLIED. All performance data presented are within the applicable limits of the helicopter.

7-4.

USE OF CHARTS.

a. Chart Explanation. The first page of each section describes the chart(s) and explains its uses. b. Coding. Chart codes are used as follows: (1) Normally dashed lines are used for example guidelines. (2) Bold lines are used for limit lines. (3) Shaded area is used for precautionary or time-limited operation. c. Reading the Charts. The primary use of each chart is given in an example and a sashed guideline is provided to help you follow the route through the chart. The use of a straight edge (ruler or page edge) and a hard fine point pencil is recommended to avoid cumulative errors. The majority of the charts provide a standard pattern for use as follows: enter first variable on top left scale, move right to the second variable, reflect down at right angles to the third

7-4

NOTE An example of an auxiliary use of the charts referenced above is as follows: Although the hover chart is primarily arranged to find torque required to hover, by entering torque available as required, maximum skid height for hover can also be found. In general, any single variable can be found if all others are known. Also, the tradeoffs between two variables can be found. For example, at a given density altitude and pressure altitude, you can find the maximum gross weight capability as free air temperature changes.

LIMITS.

Applicable limits are shown on the charts as bolt lines. Performance generally deteriorates rapidly beyond limits. If limits are exceeded, minimize the amount and time. Enter the maximum value and time above limits on DA Form 2408-13 so proper maintenance action can be taken.

7-5.

variable, reflect left and right angles to the fourth variable, reflect down, etc. until the final variable is read out at the final scale. In addition to the primary use, other uses of each chart are explained in the text accompanying each set of performance charts.

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7-6.

DATA BASIS.

The data provided generally is based on one of four categories: a. Flight Test Data. Data obtained by flight test of the aircraft by experienced flight test personnel at precise conditions using sensitive calibrated instruments. b. Derived From Flight Test. Flight test data obtained on a similar rather than the same aircraft and series. Generally small corrections will have been made. c. Calculated Data. Data based on tests, but not on flight test of the complete aircraft. d. Estimated Data. Data based on estimates using aerodynamic theory or other means but not verified by flight test.

7-7.

SPECIFIC CONDITIONS.

The data presented are accurate only for specific conditions listed under the title of each chart. Variables for which data are not presented, but which may affect that phase of performance, are discussed in the text. Where data are available or reasonable estimates can be made, the amount that each variable affects performance will be given.

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7-8.

data are not provided, thereby increasing the accuracy of performance predictions.

GENERAL CONDITIONS.

In addition to the specific conditions, the following general conditions are applicable to the performance data. a. Rigging. All airframe and engine controls are assumed to be rigged within allowable tolerances. b. Pilot Technique. Normal pilot technique is assumed. Control movements should be smooth and continuous. c. Helicopter Variation. Variations in performance between individual helicopters are known to exist; however, they are considered to be small and cannot be individually accounted for. d. Instrument Variation. The data shown in the performance charts do not account for instrument inaccuracies or malfunctions.

7-9.

PERFORMANCE DISCREPANCIES.

Regular use of this chapter will allow you to monitor instruments and other helicopter systems for malfunction, by comparing actual performance with planned performance. Knowledge will also be gained concerning the effects of variables for which

7-10. DEFINITIONS OF ABBREVIATIONS. a. Unless otherwise indicated in the following lest of abbreviations, abbreviations and symbols used in this manual conform to those established in Military Standard MIL-STD-12, which is periodically revised to reflect current changes in abbreviations usage. Accordingly, it may be noted that certain previously established definitions have been replaced by more current abbreviations and symbols. b. Capitalization and punctuation of abbreviations varies, depending upon the context in which they are used. In general, lower case abbreviations are used in text material, whereas abbreviations used in charts and illustrations appear in full capital letters. Periods do not usually follow abbreviations; however, periods are used with abbreviations that could be mistaken for whole words if the period were omitted. c. The following list provide definitions for abbreviation applies for either singular or plural applications.

LIST OF ABBREVIATIONS Abbreviation AGL ALT AVA C CAS CL CON END F FAT FLT FT FT/MIN FWD ∆F GAL

Definition Above ground level Altitude AvailableIL Celsius Calibrated airspeed Centerline ContinuousT Endurance Fahrenheit Free Air Temperature Flight Foot Feet per minute Forward Increment of equivalent flat plate drag area Gallon

Abbreviation GAL/HR GRWT GW HP HR IAS IGE IN IN HG IR KCAS KIAS KN KTAS LB LB/HR LIM

Definition Gallons per hour Gross weight Gross weight Horsepower Hour Indicated airspeed In ground effect Inch Inches of mercury Infrared Knots calibrated airspeed Knots indicated airspeed Knot Knots true airspeed Pound Pounds per hour Limit

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7-6

Definition

MAX MIN MIN MM NM NO. ∞ OGE PSI PRESS

Maximum Minimum Minute Millimeter Nautical Number Degree Out of ground effect Pounds per square inch Pressure

R/C R/D RPM

Rate of climb Rate of descent Revolutions per minute.

Abbreviation SHP SPE STA SQ F TAS TRQ USAASTA VDC VNE

XMSN

Definition Shaft horsepower SpecificationsC Station Square feetT True airspeedmile Torque United States Army Aviation Systems Test Activity Volts, direct current Velocity, never exceed (airspeed limitation) Transmission

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Section II. TORQUE AVAILABLE 7-11. DESCRIPTION.

7-12. USE OF CHARTS.

The torque available charts show the effects of altitude and temperature on engine torque. Figure 7-1 shows torque available data for 5 minute operation in terms of the allowable torque as recorded by the torque meter (%).

The primary use of the charts is illustrated by the examples. In general, to determine the maximum torque available, it is necessary to know the pressure altitude, temperature, and the aircraft calibration factor. By entering the upper left side of the chart at the known temperature (FAT -∞C), moving right to the known pressure altitude-feet, then straight down to read calibrated torque.

The power output capability of the T53-L703 engine can exceed the transmission structural limit (100%) under certain conditions. a. Figure 7-1 is applicable for maximum power engine deice off, and ECS off, 5 minute operation. b. Prolonged IGE hover may increase engine inlet temperature as much as 10 degrees Celsius; therefore, a higher FAT must be used to correct for the increase under this condition.

7-13. CONDITIONS. Charts (figure 7-1) are based upon speeds at 100 percent rpm with grade JP-4 fuel. The use of aviation jet fuel will not influence engine power, because JP-4, JP-5 and JP-8 have the same energy value per pound. Fuel grade of JP-5 or JP-8 will yield the same nautical miles per pound of fuel and being 6.8 pounds per gallon will only result in increased fuel weight per gallon.

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Figure 7-1. Maximum torque (5 minute operation) chart

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Section III. HOVER 7-14. DESCRIPTION. The hover charts (figures 7-1A thru 7-3E) show the hover ceiling and the torque required to hover at various pressure altitudes, ambient temperatures, gross weights, and skid heights. Maximum skid height for hover can also be obtained by using the torque available from figure 7-2.

7-15. USE OF CHARTS. The primary use of the hover charts is illustrated by examples. In general, to determine the hover ceiling or the toque required to hover, it is necessary to know the pressure altitude, temperature, gross weight and the desired skid height. In addition to the primary use, the hover charts can also be used to determine the predicted maximum hover height, which is needed for use of the takeoff chart (figure 7-5).

(Figure 7-2); then increase the hover gross weight by 140 pounds. Use this power available and gross weight in the hover power required chart (Figure 73A). (3) To determine maximum gross weight, first subtract 2 percent from power available (Figure 7-2); then decrease the hover gross weight determined from the hover power chart (Figure 7-3A) by 140 pounds. With the rotor blade erosion protection coating and polyurethane tape installed, it will necessary to make the following corrections. Add 2 percent to the hover torque required, for OGE and IGE, as determine from (Figure 7-3A). In (Figure 7-3), subtract 100 pounds from the maximum gross weight to hover. When determining maximum hover wheel weight, enter the chart at the gross weight plus 100 pounds.

7-16. CONDITIONS. 7-16A. HOVER CEILING. a. The hover charts do not account for the effect of an IR suppressor device. The hover ceiling chart (Figure 7-3) is not usable if a suppressor device is installed. The IR Scoop Suppressor creates a download of approximately 140 pounds. b. For the IR Scoop Suppressor: (1) To determine hover torque required, enter the hover power required chart (Figure 7-3A) at a gross weight of 140 pounds heavier than the actual gross weight. (2) To determine predicted maximum hover height, first subtract 2 percent from power available

In ground effect (IGE) and out ground effect (OGE) hover ceiling charts (figures 7-1A thru 7-3E) are based upon engine manufacturer's minimum specification power for T53-L-703 engine with installation losses. Those charts reflect maximum hovering capability of helicopter in zero wind conditions, whereas hover performance shown in Basic Flight Manual is reduced to ensure adequate tall rotor control margins in relative winds up to 20 knots from any direction. Caution, therefore, should be exercised when hovering at high GW and high HD, as tail rotor control margins may not be available, particularly when winds are within critical relative wind azimuth area (figure 7-3F).

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TBD

Figure 7-1A. Hover torque required (Sheet 1 of 2)

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TBD

Figure 7-1A. Hover torque required (Sheet 2)

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Figure 7-2. Hover ceiling in ground effect (IGE) (5-Minute Operation) - Heater On

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Figure 7-2A. Hover ceiling in ground effect (IGE) (5-Minute Operation) - Heater Off

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Figure 7-3. Hover ceiling in ground effect (IGE) (Continuous Operation) - Heater On

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Figure 7-3A. Hover ceiling in ground effect (IGE) (Continuous Operation) - Heater Off

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Figure 7-3B. Hover ceiling out of ground effect (OGE) (5 Minute Operation) - Heater On

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Figure 7-3C. Hover ceiling out of ground effect (OGE) (5 Minute Operation) - Heater Off

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Figure 7-3D. Hover ceiling out of ground effect (OGE) Continuous Operation) - Heater On

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Figure 7-3E. Hover ceiling out of ground effect (OGE) (Continuous Operation) - Heater Off

7-14D

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Figure 7-14F. Critical relative wind azimuths for hover flight

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Section IV. CONTROL MARGIN 7-17. CONTROL MARGIN CHARTS. Sheet 1 of the control margin chart (figures 7-4) shows the maximum right crosswind in which directional control can be maintained as a function of pressure altitude, temperature, and gross weight. Sheet 2 of the control margin chart (figures 7-4) shows the combinations of relative wind velocity and azimuth which may result in marginal directional or longitudinal control. Use of the control margin chart is illustrated by the example on sheet 1. Ten percent pedal margin (full righ t to full left) is cons id ered adequate f or directional control when hovering. The shaded area on sheet 1 indicates conditions where the directional control margin may be less than ten percent in zero

wind hover. The shaded area on sheet 2 labeled DIRECTIONAL, indicates conditions where the directional control margin may be less than ten percent for crosswind components in excess of those determined from sheet 1. The shaded area on sheet 2 labeled LONGITUDINAL indicates wind conditions where longitudinal cyclic control margin may be less than 10 percent. These charts are based on control margins only.

7-18. CONDITIONS. The control margin charts are based on test of inground-effect (IGE) translational flight over a level surface at 324 rotor 6600 engine rpm. Use of these charts is to determine if adequate control margin will be available for IGE and OGE hover in winds or low speed translational flight.

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Figure 7-4. Control margin chart (Sheet 1 of 2)

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Figure 7-4. Control margin chart (Sheet 2 of 2)

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Section V. TAKEOFF capability is limited, this technique results in increased distance to clear an obstacle.

7-19. DESCRIPTION. The takeoff chart (figure 7-5) shows the distance to clear various obstacle height, based upon hover height capabilities. the upper chart grid presents data from climbout at a constant INDICATED airsp eed . The tw o lo wer g rid s p re sent d ata f or climbouts at various TRUE airspeeds. Figure 7-5, sheet 1, is based upon level acceleration at a three foot skid height, sheet 2 is based upon a climb and acceleration from a three foot skid height and sheet 3 is based upon a climb and acceleration from a 15 foot skid height. a brief explanation of the techniques follows:

c. Level Acceleration from a 15 Foot Skid Height. Level acceleration from a 15 foot skid height hover (figure 7-5, sheet 3), is used for loads suspended beneath the helicopter. The takeoff is accomplished by applying full power while accelerating though translational lift. Climbout airspeed is maintained at 5 to 10 knots above translational lift speed when the minimum distance over a 50-foot obstacle is required. This takeoff technique is recommended for normal to heavy weight takeoffs when excess power available is limits.

NOTE These distances are obtained with maximum torque available (5 minute operation). a. Level Acceleration from a Three Foot Skid Height. Level acceleration from a three foot skid height hover (figure 7-5), sheet 1) yields successful takeoffs when the power available is sufficient to hover at skid heights above three feet. The takeoff is accomplished by applying full power while accelerating through translational lift at a skid height of three feet. Climbout airspeed is maintained at 5 to 10 knots above translational lift speed when the minimum distance over a 50-foot obstacle is required. this takeoff technique is recommended for normal, heavy weight takeoffs when excess power available is limited. b. Climb and Accelerate from the Three Foot Skid Height. Climb and accelerate from a three foot skid height lover (figure 7-5, sheet 2) yields the best takeoff performance at light weights. The power available has to be sufficient to climb and accelerate simultaneously. Its skid height increases, the ground effect augmentation decreases, and the helicopter requires more power to continue the climb. When sufficient power is not available, it will stop climbing and gradually settle. Before resuming the climb, altitude must be traded for airspeed. As airspeed increases, the power required for level flight decreases and sufficient excess power becomes available to continue the climb. When hovering

The standard airspeed system is inaccurate in the climbout airspeed range requ ired for maximum performance (minimum distance) takeoffs. Airspeed is maintained by reference to the helicopter's attitude. Airspeeds required for maximum performance places the helicopter in the "avoid" region of the height-velocity chart and should be used only when maximum performance is required. NOTE The hover heights shown on the chart are only a measure of the aircraft's climb capability and do not imply that a higher than normal hover height should be used during the actual takeoff.

7-20. USE OF CHARTS. The primary use of these charts are illustrated by the charts examples. The main consideration for takeoff performance is the hovering skid height capability, which includes the effects of pressure altitude, free air temperature, gross weight, and torque. Hover height capability is determined by use of the hover chart, figure 7-2. A hover check can be made to verify the hover capability. If winds are present, the hover check may disclose that the helicopter can actually hover at a greater skid height than the calculated

Rev. 7

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value, since the hover chart is based upon clam wind conditions.

7-21. CONDITIONS. a. Wind. The takeoff chart is based upon calm wind conditions. Since surface wind velocity and direction cannot be accurately predicted, all takeoff planning should be based upon calm wind conditions. Takeoff into any prevailing wind will improve the takeoff performance.

7-20

A tailwind during takeoff and climbout will increase the obstacle clearance distance and could prevent a successful takeoff. b. Power Settings. All takeoff performance data are based upon the torque used in determining the hover capabilities in figure 7-3.

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Figure 7-5. Takeoff chart (Sheet 1 of 3)

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Figure 7-5. Takeoff chart (Sheet 2 of 3)

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Figure 7-5. Takeoff chart (Sheet 3 of 3)

7-23/(7-24 blank)

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Section VI. CRUISE 7-22. CRUISE. The cruise charts (figure 7-6, sheets 1 through 48) show the torque pressure and fuel flow required for level flight at various pressure altitudes, airspeed, and gross weights. The cruise charts are presented for a clean drag configuration. However, any desired drag configuration (figure 7-7 authorized configurations) can be solved by using the difference between the drag for the desired configuration and the clean configuration. This solution is ratioed to the drag for 5 square foot equivalent flat plate drag area. Add or subtract the calculated ▲ % if the desired configuration to the clean configuration to solve the total torque required for the desired configuration. Then read the fuel flow for the desired configuration. Thus, the effects of different drag resulting from changing configurations can be determined by applying the ▲F-SQ FT from the Drag Chart (figure 7-7) to the cruise charts (figure 7-6).

7-23. USE OF CHARTS.

Cruise flight is restricted to 319 to 324 Rotor RPM (6500 to 6600 Engine RPM.) Cruise at 324 Rotor/6600 Engine RPM is recommended. The cruise chart data for true airspeeds above 40 KTAS is based on 314 Rotor/6400 Engine RPM. Until the cruise charts are revised performance planning shall be accomplished using the procedures and torque corrections from Figure 7-6. The primary use of the charts is illustrated by the examples provided in figure 7-6. The first step for chart use is to select the proper chart, pressure altitude and anticipated free air temperature, (paragraph 7-2). Normally, sufficient accuracy can be obtained by selecting the chart nearest to the planned cruising altitude and FAT, or the next higher altitude and FAT. If greater accuracy is required, interpolation between altitude and/or temperatures will be required (figure 7-7, sheet 1, example A). You may enter the charts on any side: TAS, IAS, torque pressure, or fuel flow, and then move vertically or horizontally to the gross weight, then to the other three parameters. Maximum performance conditions are determined by entering the chart where the maximum range of maximum endurance and rate of climb lines intersect the appropriate gross weight; then read airspeed, fuel

flow, and torque pressure. For conservatism, use the gross weight at the beginning of cruise flight. For greater accuracy on long flights it is preferable to determine cruise information for several flight segments in order to allow for decreasing fuel weight (reduced gross weight). The following parameters contained in each chart are further explained as follows: a. Airspeed. True and indicated airspeeds are presented at adjacent sides of each chart. On any chart, indicated airspeed can be directly converted to true airspeed (or vice versa) by reading IAS vs TAS line. Maximum permissible airspeed (VNE) limits appear on some charts. If no line appears. VNE is above the limits of the chart. b. Torque Pressure (%). Since pressure altitude and temperature are fixed for each chart, torque % required varies according to gross weight and airspeed. c. Fuel Flow. Fuel flow is shown as a function of torque %. On any chart, torque % may be converted directly to fuel flow without regard for other chart information. All fuel flow information is presented with bleed air heater and de-ice off. Add two percent fuel flow (about 14 lb/hr) for bleed air heater on and increase fuel flow three percent (approximately 21 lb/ hr) for anti-ice on. If both are operating, add five percent fuel flow (about 35 lb/hr) to chart values. d. Maximum Range. The maximum range lines indicate the combinations of weight and airspeed that will produce the greatest flight range per gallon of fuel under zero wind conditions. When a maximum range condition does not appear on a chart it is because the maximum range speed is beyond the maximum permissible speed (VNE); in such cases, use (VNE) cruising speed to obtain maximum range. e. Maximum Endurance and Rate of Climb. The maximum endurance and rate of climb lines indicate the airspeed for minimum torque pressure required to maintain level flight for each gross weight, FAT and pressure altitude. Since minimum torque pressure will provide minimum fuel flow, maximum flight endurance will be obtained at the airspeed indicated.

7-24. CONDITIONS. The cruise charts are based on op eration of 324/rotor/6600 engine rpm and bleed air, heater, and anti-ice off. Rev. 7

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Figure 7-6. Cruise chart, clean configuration, –30°C, sea level, (Sheet 1 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 2000 feet, Sheet 2 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 4000 feet, (Sheet 3 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 6000 feet, (Sheet 4 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 8000 feet, (Sheet 5 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 10000 feet, (Sheet 6 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 12000 feet, (Sheet 7 of 48)

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Figure 7-6. Cruise chart, clean configuration, –30°C, 14000 feet, (Sheet 8 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, sea level, (Sheet 9 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 2000 feet, (Sheet 10 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 4000 feet, (Sheet 11 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 6000 feet, (Sheet 12 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 8000 feet, (Sheet 13 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 10000 feet, (Sheet 14 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 12000 feet, (Sheet 15 of 48)

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Figure 7-6. Cruise chart, clean configuration, –15°C, 14000 feet, (Sheet 16 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, sea level, (Sheet 17 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 2000 feet, (Sheet 18 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 4000 feet, (Sheet 19 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 6000 feet, (Sheet 20 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 8000 feet, (Sheet 21 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 10000 feet, (Sheet 22 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 12000 feet, (Sheet 23 of 48)

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Figure 7-6. Cruise chart, clean configuration, –0°C, 14000 feet, (Sheet 24 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, sea level, (Sheet 25 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 2000 feet, (Sheet 26 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 4000 feet, (Sheet 27 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 6000 feet, (Sheet 28 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 8000 feet, (Sheet 29 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 10000 feet, (Sheet 30 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 12000 feet, (Sheet 31 of 48)

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Figure 7-6. Cruise chart, clean configuration, +15°C, 14000 feet, (Sheet 32 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, sea level, (Sheet 33 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 2000 feet, (Sheet 34 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 4000 feet, (Sheet 35 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 6000 feet, (Sheet 36 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 8000 feet, (Sheet 37 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 10000 feet, (Sheet 38 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 12000 feet, (Sheet 39 of 48)

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Figure 7-6. Cruise chart, clean configuration, 30°C, 14000 feet, (Sheet 40 of 48)

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Figure 7-6. Cruise chart, clean configuration , 45°C, sea level, (Sheet 41 of 48)

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Figure 7-6. Cruise chart, clean configuration, 45°C, 2000 feet, (Sheet 42 of 48)

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Figure 7-6. Cruise chart, clean configuration, 45°C, 4000 feet, (Sheet 43 of 48)

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Figure 7-6. Cruise chart, clean configuration, 45°C, 6000 feet, (Sheet 44 of 48)

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Figure 7-6. Cruise chart, clean configuration, 45°C, 8000 feet, (Sheet 45 of 48)

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Figure 7-6. Cruise chart, clean configuration, 55°C, sea level, (Sheet 46 of 48)

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Figure 7-6. Cruise chart, clean configuration, 55°C, 2000 feet, (Sheet 47 of 48)

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Figure 7-6. Cruise chart, clean configuration, 55°C, 4000 feet, (Sheet 48 of 48)

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Section VII. DRAG

The drag chart (figure 7-7, sheet 1 of 2) shows the equivalent flat plate drag area changes for additional authorized configurations. there is no increase in drag with cargo doors fully open. The upper left portion of figure 7-7, sheet 2 of 2, presents drag areas of typical external loads as a function of the load frontal area. The balance of the chart shows the additional toque required in level flight due to the increase in drag caused by external loads or aircraft modifications or authorized configurations. The IR Scoup Suppressor has a drag of two square feet.

is necessary to know the drag area change, true airspeed, pressure altitude, an free air temperature. Enter at known drag area change, move right to TAS move down to pressure altitude, move left to FAT, then move down and read change in torque. In addition, by entering the chart in the opposite direction, drag area change may be found from a known torque change. this chart is used to adjust cruise chart torque and fuel flow due to equivalent flat plate drag area change (▲F). For frontal areas exceeding values shown on figure 7-7 (sheet 2 of 2) use a smaller value and multiply, e.g. 36 sq. ft. = 12 sq. ft. X3.

7-26. USE OF CHART.

7-27. CONDITIONS.

The primary use of the drag chart is illustrated by the example. To determine the change in torque, it

The drag chart is based upon 324 rotor/6600 engine rpm.

7-25. DESCRIPTION.

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Figure 7-7. Drag chart (Sheet 1 of 2)

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Figure 7-7. Drag chart (Sheet 2 of 2)

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Section VIII. CLIMB PERFORMANCE/CLIMB — DESCENT 7-28. DESCRIPTION — CLIMB PERFORMANCE. Figure 7-8 represents a synthesis of the cruise charts to ease estimation of the climb portion of the flight plan. The chart shows the time, distance, and fuel required to climb from an initial altitude to a final altitude. The chart provides for variation in gross weight and ambient temperature and may be used for minor configuration deviations.

7-29. USE OF CHART — CLIMB PERFORMANCE. The primary uses of the climb performance charts are illustrated by the chart examples.

7-30. CONDITIONS — CLIMB PERFORMANCE. The climb-performance chart represents climb at optimum conditions, that is, at the best rate-ofclimb airspeed and at maximum power available (for 30-minute operation). Climb is assumed to be at 60 knots indicated airspeed as this is near the airspeed for maximum rate-of-climb at most atmospheric conditions. Warmup and taxi fuel are not included in fuel calculations. Climb performance is calculated for 324 rotor/6600 engine rpm. The charts are based upon a no-wind condition; therefore, distance traveled will not be valid when winds are present.

7-31. DESCRIPTION — CLIMB-DESCENT. Figure 7-9, shows the change in torque (above or below torque required for level flight under the same gross weight and atmospheric conditions) to obtain a given rate of climb or descent.

7-32. USE OF CHART — CLIMB-DESCENT. The primary uses of the climb-descent chart are illustrated by the chart examples. a. Torque Change. The torque change obtained from the grid scale must be added to (for climb) or subtracted from (for descent) the torque required for level flight to obtain a total climb or descent torque. (Torque required for level flight is obtained from the appropriate cruise chart.) b. Known Torque Change. By entering the bottom of the grid with a known torque change, moving upward to the gross weight, then left, the corresponding rate of climb or descent may also be obtained.

7-33. CONDITIONS — CLIMB-DESCENT. The climb-descent chart is based on the use of constant rotor or engine rpm. A decrease in rpm could decrease the rate of climb or increase the rate of descent shown.

Rev. 1

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Figure 7-8. Climb performance

Rev. 6

7-85

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Figure 7-9. Climb — descent chart

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Section IX. FUEL FLOW 7-34. DESCRIPTION. The idle fuel flow chart (figure 7-10) shows the fuel flow at engine idle. The fuel flow vs torque chart (figure 7-11) shows fuel flow at 324 rotor/6600 engine rpm in poundsper-hour versus percent torque for pressure altitudes from sea level to 20,000 feet and for 0°C free air temperature.

7-35. USE OF CHART. The primary use of the idle fuel flow chart is illustrated by the example. To determine the idle fuel flow, it is necessary to know the pressure altitude and free air temperature. Enter at the pressure altitude, move right to FAT in appropriate grid, then move down and read fuel flow on the bottom scale.

The primary use of the fuel flow vs torque chart is illustrated by the example. To determine fuel flow, it is necessary to know the torquemeter percent (%Q) and the FAT as well as the pressure altitude. Fuel flow will increase about 2 percent with the bleed air heater on and 3 percent with de-ice on. When both systems are on, fuel flow will increase 5 percent. Also a range of endurance penalty should be accounted for when working cruise chart data. A fairly accurate rule of thumb to correct fuel flow for temperatures other than 0°C FAT is to increase/ decrease fuel flow 1 percent for each 10°C increase/ decrease in FAT.

7-36. CONDITIONS. The fuel flow charts are based upon the use of JP-4 fuel. The change in fuel flow when using other jet fuels is insignificant.

Rev. 7

7-87

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Figure 7-10. Idle fuel flow

7-88

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Figure 7-11. Fuel flow vs torque, 0°C

Rev. 1

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Section X. NAUTICAL MILES AND TIME AND RANGE 7-37. DESCRIPTION. The nautical miles per pound of fuel chart (figure 712) shows the nautical miles per pound of fuel that the helicopter will use during a given mission. The only information required is the fuel flow in pounds per hour and the cruise airspeed. The time and range vs fuel chart (figure 7-13) shows the enroute time and distance that the helicopter can cover while in level flight with calm winds. The only information required is the cruise fuel, the fuel

7-90

flow in pounds per hour and the cruise true airspeed. Calculations are based on zero fuel reserve.

7-38. USE OF CHARTS. See example below and on page 7-91.

7-39. CONDITIONS. The nautical miles per pound of fuel chat and the time and range chart is based on calculated data.

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Figure 7-12. Nautical miles per pound of fuel

7-91

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7-92

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Figure 7-13. Time and range versus fuel

7-93

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Section XI. AUTOROTATIONAL GLIDE CHARACTERISTICS 7-40. AUTOROTATIONAL GLIDE CHARACTERISTICS. The Autorotational GLIDE Characteristic chart (figure 7-14) shows both the rate of descent and the glide ratio (horizontal distance covered divided by the altitude lost in a fixed time) for various rotor speeds and indicated airspeeds that the helicopter is capable of without power. The chart shows the tradeoffs available to the pilot for his autorotational descent. When considering autorotational flight, the following points should be considered: a. Use of low rotor speed to decrease the rate of descent should be limited to 500 feet or more above the ground. b. Use of low rotor speed for maximum glide should be limited to 500 feet or more above the ground. c. Rate of descent is basically controlled by airspeed and rotor speed. d. An increase in rotor speed will provide a slight increase in rate of descent and large increase in rotor energy and lift that is needed during the flare and landing. e. Effectiveness of the cyclic flare is dependent upon: (1) Entry rotor RPM. (2) Entry airspeed. (3) Rate and steepness of the flare.

(2) Gross weight. (3) Density altitude. (4) Rate of descent at application.

NOTE If cannot be emphasized enough that high rotor speeds are required for successful cyclic flares and landing maneuvers during autorotation. This high rotor speed must be regained prior to the landing portion. The speeds for maximum glide distance is 80 to 85 KIAS. The speeds for minimum rate of descent or maximum time to descend is 45-50 KIAS. Example: Find the maximum glide distance, rate of descent and associated airspeed for 5500 feet above ground level. Solution: a. The maximum glide distance corresponds to the best glide ration. The chart (figure 7-14) on the bottom, shows that 85 KIAS is the speed for the highest glide ratio, which is 5.32 at 94% NR. b. Proceed to the upper portion of the chart and read a corresponding rate of descent of 1820 feet per minute.

(4) Gross weight. (5) Density altitude. (6) Rate of descent. f. Effectiveness of collective pitch for landing is dependent upon: (1) Rotor RPM at application.

7-94

c. Multiply the height above ground (with the last 500 feet not included because it is to be reserved for regaining rotor speed prior to the flare) by the glide ratio found in step a to get 5.32 x 5000 feet, or 26,600 feet. Divide it by 6076 feet per nautical mile to get 4.4 nautical miles. this is the maximum horizontal distance the helicopter can glide at 85 KIAS, 94% NR and 1820 feet per minute rate of descent.

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Figure 7-14. Autorotational glide characteristics chart

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Section XII. AIRSPEED LIMITS 7-41. DESCRIPTION. The airspeed operating limits chart (figure 7-15) shows the level flight indicated airspeed limits as a function of density altitude and gross weight. With a known pressure altitude and FAT, the indicated airspeed limit (Vne) can be obtained as shown in the Example. The top part of the chart can also be used to read a density altitude for a given pressure altitude and FAT.

7-41A. Vne is 80 knots at or below 10,000 feet Hp for all GW with external cargo on suspension unit.

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Rev. 6

Decrease Vne 3 knots per thousand feet above 10,000 feet HD.

Airspeed with external cargo is limited by c o n t r o l l a b i l i t y. C a u t i o n s h o u l d b e exercised when carrying cargo as handling characteristics may be affected due to size, weight, and shape of cargo load.

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Figure 7-15. Airspeed operating limits

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Section XIII. AIRSPEED CORRECTION 7-42. DESCRIPTION. Airspeed correction chart (figure 7-16) shows the calibrated airspeed as a function of indicated airspeed for level flight condition only. Indicated airspeed is directly converted to calibrated airspeed by the use of this chart.

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Figure 7-16. Airspeed correction chart

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Section XIV. TEMPERATURE 7-43. DESCRIPTION. The temperature conversion chart (figure 7-17) can be used to convert degrees F to degrees C (or vice versa), within a range of -60 ∞C to +60 ∞C or -76 ∞F to +140 ∞F.

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Rev. 1

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Figure 7-17. Temperature conversion chart

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Section XV. PERFORMANCE PLANNING 7-44. PERFORMANCE PLANNING CARD. This card (figure 7-18) is provided to assist you in recording data applicable to the mission and may be reproduced at the local level.

7-45. PERFORMANCE PLANNING SEQUENCE. The following sequence is provided to aid you in preparing the performance planning card. a. Pressure Altitudes. Obtain departure point pressure altitude by setting 29.92 in. hg at the altimeter barometric pressure scale and reading pressure altitude from outer scale pointers. Record in space provided under departure heading. Estimate pressure altitude increase above departure elevation; (if destination is below departure elevation, subtract difference in elevation). Record pressure altitudes in space provided under climb, cruise, and arrival headings. b. Free Air Temperature. Obtain the local fee air temperature and record in space provided under departure heading. Estimate FAT for climb, cruise, and arrival by subtracting 2 degrees C for each 1000 feet altitude increase above departure point; (if destination is below departure elevation, add 2∞C for each 1000 feet difference in elevation). Record temperature in space provided under climb, cruise and arrival headings.

(4) Determine and record torque required to hover at desired skid height. See figure 7-2. (5) Determine if takeoff to hover can be made by comparing torque required for desired skid height with maximum torque available. For takeoff, torque available must be greater than torque required. b. Obstacle clearance. (1) Determine and record obstacle height. (2) Determine and record distance to obstacle. (3) Select and record the airspeed that will allow the helicopter to safely clear obstacle. See figure 7-5. c. Climb. (1) Conservatively, using departure gross weight, determine and record speed for maximum rate of climb, IA and the torque pressure for cruise at this speed. See figure 7-6 (sheets 1 through 48) using pressure altitude and FAT previously determined. (2) Record maximum torque pressure available and maximum fuel flow. See figure 7-6 (sheets 1 through 48). (3) Subtract level flight torque required from maximum torque available to obtain change in excess torque for climb. Record this value. (4) Determine rate of climb from figure 7-8.

NOTE If required, see figure 7-17 for temperature conversion chart (Fahrenheit to Celsius, or vice versa). c. Departure. (1) Calculate and record the departure gross weight. (2) Determine and record torque available. See figure 7-1.

d. Cruise. (1) Select and record cruise speed (TAS). (2) Calculate and record gross weight at beginning of cruise segment or average weight during cruise segment. (3) Using pressure altitude and FAT previously determined from figure 7-6 (sheets 1 through 48) determined and record the following: (a) Cruise speed (IAS). (b) Cruise torque pressure.

(3) Determine and record torque required at five feet.

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Rev. 5

(c)

Cruise fuel flow.

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e. Arrival. (1) Calculate and record gross weight. (2) Select and record approach airspeed (IAS). (3) Using the method described for departure in steps c. (2), (3), and (4) preceding, determine whether a hovering approach to landing can be accomplished. Torque available for landing must exceed torque required. See figures 7-1, and 7-5.

NOTE Performance information obtained may make it necessary to alter gross weight, airspeed, altitude, or other variables in order to safely operate the helicopter. If any of these variables are changed on one chart, corresponding changes will be necessary on all other charts where that information is used.

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Figure 7-18. Performance planning card

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Rev. 5

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CHAPTER 8 NORMAL PROCEDURES Section I. MISSION PLANNING 8-1.

MISSION PLANNING.

Mission planning begins when the mission is assigned and extends to the preflight check of the helicopter. It includes, but is not limited to checks of operating limits and restrictions; weight balance and loading; performance; publications; flight plan and crew and passenger briefing. The pilot in command shall ensure compliance with the contents of this manual that are applicable to the mission. NOTE Instruct ground personnel to discharge helicopter static electricity, before attaching cargo sling, by touching airframe with a ground wire. If a metal sling is used, hook up ring can be struck against cargo hook to discharge static electricity. If contact has been lost after initial grounding, helicopter should be electrically regrounded and, if possible, contact maintained until hook up is complete.

8-2. OPERATING LIMITS AND RESTRICTIONS. The minimum, maximum, normal and cautionary operational ranges represent careful aerodynamic and structural calculation, substantiated by flight test data. These limitations shall be adhered to during all phases of the mission. Refer to Chapter 5, OPERATING LIMITS AND RESTRICTIONS, for detailed information.

8-3.

WEIGHT, BALANCE, AND LOADING.

The helicopter shall be loaded, cargo and passengers secured, and weight and balance verified in accordance with Chapter 6, WEIGHT, BALANCE

AND LOADING. This helicopter is in weight and balance Class I, and requires a weight and balance clearance only when loaded in other than a normal manner. The helicopter weight and center-of-gravity conditions shall be within the limits prescribed in Chapter 5, OPERATING LIMITS AND RESTRICTIONS.

8-4.

PERFORMANCE.

Refer to Chapter 7, PERFORMANCE DATA, to determine the capability of the helicopter for the entire mission. Consideration shall be given to changes in performance resulting from variation in loads, temperatures, and pressure altitudes. Record the data on the Performance Planning Card for use in completing the flight plan and for reference throughout the mission.

8-5.

FLIGHT PLAN.

A flight plan shall be completed and filed in accordance with local regulations.

8-6.

CREW AND PASSENGER BRIEFING.

A crew briefing shall be conducted to ensure thorough understanding of individual and team responsibilities. The briefing should include, but not be limited to, copilot, crew chief, mission equipment operator, and ground crew responsibilities and the coordination necessary to complete the mission in the most efficient manner. A review of visual signal is desirable when ground guides do not have direct voice communications link with the crew. Refer to Section VI for passenger briefing. See figure 8-1 for typical danger areas.

Section II. OPERATING PROCEDURES AND MANEUVERS basic flight principles are avoided. Only the duties 8-7. OPERATING PROCEDURES AND of the minimum crew necessary for the actual MANEUVERS. This section deals with normal procedures, and includes all steps necessary to ensure safe and efficient operation of the helicopter from the time a preflight begins until the flight is completed and the helicopter is parked and secured. Unique feel, characteristics and reaction of the helicopter during various phases of operation and the techniques and procedures used for taxiing, takeoff, climb, etc., are described including precautions to be observed. Your flying experience is recognized; therefore,

operation of the helicopter are included.

Operation of the helicopter with no load on the external cargo suspension hook is authorized under the standard airworthiness certificate under VFR or IFR conditions without removing the unit from the helicopter. The installation and use of the rear view mirror contained in the kit is left to operators discretion.

Rev. 6

8-1

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Figure 8-1. Danger area

8-2

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The rear view mirror shall be covered or removed for night flight. NOTE VFR OPERATION. With a load attached to the suspension assembly, operation shall be conducted in accordance with appropriate operating rules for external loads under VFR conditions.

IFR OPERATION. External load operations are permitted provided the operator substantiates to the Administrator that the rotorcraft ó load combinations meets IFR handling requirements and insures that the Rotorcraft External Load Operator Certificate reflects same with appropriate restrictions.

Rev. 6

8-2A/(8-2B blank)

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8-8.

ADDITIONAL CREW DUTIES.

Additional crew duties are covered as necessary in Section VI, CREW DUTIES. Mission equipment checks are contained in Chapter 4, MISSION EQUIPMENT. Descriptions of functions, operations, and effect of controls are covered in Section IV, FLIGHT CHARACTERISTICS, and are repeated in this section only when required for emphasis. Checks that must be performed under adverse environmental conditions, such as desert and cold weather operations, supplement normal procedures checks in this section and are covered in Section V, ADVERSE ENVIRONMENTAL CONDITIONS.

8-9.

CHECKLIST.

Normal procedures are given primarily in checklist form, and amplified as necessary in accompanying paragraph form, when a detailed description of a procedure or maneuver is required. A condensed version of the amplified checklist, omitting all explanatory text, is contained in the Operator and Crewmember Checklist, BHT PUB 92-004-CL. To provide for easier cross-referencing, the procedural steps are numbered to coincide with the corresponding numbered steps in the checklist.

8-10. CHECKS. The checklist includes items for all flight conditions with annotative indicators immediately preceding the check to which they are pertinent. The symbol ★ preceding steps of the checklist indicates that detailed procedures for those checks are included in the performance checks section located at the back of the condensed checklist. When a helicopter is flown by the same flight crew on a mission requiring intermediate stops it is not necessary to perform all of the normal checks. The steps that are essential for safe helicopter operations on intermediate stops are designated as "thru-flight" checks. An asterisk (*) indicates that performance of steps is mandatory for all "thru-flights" when there has been no change in crew. T he aste ris k app lies only to checks performed prior to takeoff. An (O) indicates a requirement if the equipment is installed.

8-11. CHECKLIST CALLOUT. Pilots and crewmembers shall not rely on memory for accomplishment of prescribed operational checks except for those immediate action emergency procedures that shall be memorized for safe

helicopter emergency operation. Oral callout and confirmation of checklist items shall be accomplished by pilot and crewmembers.

8-12. BEFORE EXTERIOR CHECK. a. Publications ó Check DA Forms 2408-12, -13, and -14; DD Form 365-4; locally required forms and publications; and availability of Operators Manual (-10), TM 1500-1 and Checklist (-CL). b. BAT switch ó ON. c. Lights, ON ó Check landing, search, anticollision, position, and interior lights for condition and operation as required; secure landing and search light as desired; then OFF. *d. Fuel ó Check quantity for first flight of the day and after refueling. Caps secure. *e. Fuel sample ó Check for contamination for first flight of the day and after refueling. If the fuel pumps, sumps, and filter have not been drained by maintenance personnel, drain as follows: (1) Pumps and sumps ó Drain and check. (2) FUEL switch ó ON. (3) Filter ó Drain and check. Check fuel lines and fittings for leaks. (4) Auxiliary fuel tank ó Drain tank and filter. Check for leaks. (5) FUEL switch ó OFF. f. Cargo hook ó Check operation, if use is anticipated, refer to Chapter 4 MISSION EQUIPMENT, for checks of the system. g. switch ó OFF. *h. Helicopter covers, tiedowns, and grounding cables ó Removed and secured, except aft main rotor tiedown.

8-13. EXTERIOR CHECK. The pilot is responsible for all preflight checks to determine whether the helicopter is in condition for

Rev. 1

8-3

BHT PUB-92-004-10 safe flight. For suggested preflight exterior check, refer to figure 8-2. NOTE The preflight check is not intended to be a detailed mechanical inspection, but simply a guide to help the pilot check the condition of the helicopter. It may be made as comprehensive as conditions warrant at the discretion of the pilot. All areas checked shall include a visual check for evidence of corrosion, particularly when helicopter is flown near or over salt water or in areas of high industrial emissions.

8-14. AREA 1 — FUSELAGE AND MAIN ROTOR. *a. Forward main rotor blade ó Check condition. b. Cabin top ó Check condition of ventilators, windshields, wipers, FAT probe, and WSPS. c. Pilot door ó operation.

Check condition, security, and

d. Pilot seat, seat belt and shoulder harness ó Check condition and security. e. Fire extinguisher ó not broken) and security.

Check for charge (seal

8-15. AREA 2 — CABIN NOSE. a. Radio compartment ó Secure of all equipment. Check battery, if installed. Secure door.

d. Cabin doors ó and security.

Check condition, operation,

e. Navigation and position lights ó condition and security.

Check

f. Fuselage ó Check condition.

8-17. AREA 4 — AFT FUSELAGE LEFT SIDE. *a. Engine compartment ó Check fuel and oil lines and connections for conditions, security, and leaks. Check general condition. Cowling secure. b. Radio and electrical compartments ó Check condition, circuit breakers in and components secure. Secure access doors. c. 2nd tail rotor driveshaft compartment ó Check. d. Fuselage ó Check condition.

8-18. AREA 5 — TAILBOOM LEFT SIDE. a. Tail rotor driveshaft access covers ó Check condition and security of tail rotor driveshaft access covers. b. Check grease packed couplings for evidence of overheating and/or grease leakage. Secure access door. c. Tailboom ó Check condition of skin. d. Synchronized elevator ó Check condition and security.

b. Cabin lower area ó Check condition of windshield, antennas, and fuselage. Check for loose objects inside which might jam controls. (O)c. Cargo suspension mirror ó Check security and cover installed. Uncover and adjust if cargo operations are anticipated.

8-16. AREA 3 — FUSELAGE LEFT SIDE. a. Landing gear ó Check condition and security of cross tubes, skid, and skid shoes. Ground handling wheels removed. b. Copilot door ó Check condition, security and operation. c. Copilot seat, seat belt and shoulder harness ó Check condition and security; secure belt and harness if seat is not used during flight.

8-4

Rev. 5

Do not rotate tail rotor by hand using the tail rotor blades. *e. Main rotor blade ó Check condition, rotate in normal direction 90∞to fuselage, tiedown removed.

8-19. AREA 6 — TAILBOOM AFT. a. Antenna ó Check condition and security. b. Tail skid ó Check condition and security. c. Positions lights ó security.

Check condition and

d. Vertical fin ó Check condition and security.

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Figure 8-2. Exterior check diagram

8-5

BHT PUB-92-004-10 *e. Tail rotor gearboxes (90∞and 42∞) ó Check general condition, oil levels, filler caps secure, and chip detector wires secure. Check grease packed couplings for evidence of overheating and/or grease leakage. *f. Tail rotor ó Check condition and free movement on flapping axis. The tail rotor blades should be checked as the main rotor is rotated. Visually check all components for security.

8-20. AREA 7 — TAILBOOM RIGHT SIDE. *a. Main rotor blade ó Check condition, rotate in normal direction 90∞to fuselage, tiedown removed. b. Synchronized elevator ó Check condition and security. c. Tailboom ó Check condition of skin. d. Baggage compartment ó Check condition.

8-21. AREA B — AFT FUSELAGE RIGHT SIDE. a. Fuselage ó Check condition of skin. b. Engine exhaust ó Check free of obstructions. c. 2nd Tail rotor driveshaft compartment ó Check d. Oil cooling fan compartment ó Check condition of fan, flight control and cables, tail rotor servo for leaks and security and battery if installed. Check for installation of structural support. Check tailboom attachment bolts. Secure door.

b. Transmission and hydraulic filler caps ó Secure. *c. Hydraulic fluid level ó Check. d. Engine driveshaft ó security.

Check condition and

e. Engine air intake ó unobstructed.

Check clean and

f. Anticollision light ó Check condition. g. Engine and transmission cowling ó condition and security. h. Position security.

lights

ó

Check

Check

condition

and

i. Antennas ó Check condition and security. j. Cabin top ventilators ó Check unobstructed. k. Pitot tube unobstructed.

ó

Check

security

and

l. WSPS ó Check.

8-24. INTERIOR CHECK — CABIN. *a. Transmission oil level ó Check. *(O)b.Cargo ó Check as required for proper loading and security.

e. Heater compartment ó Check heater for condition and security. Check area clear of obstructions. Secure door.

*c. Loose equipment ó Stow rotor blade tiedown, pitot tube cover, tailpipe cover and other equipment.

*f. Engine compartment ó Check fuel and oil lines and connections for condition, security, and leaks. Check fluid levels and general condition. Cowling secure.

(O)d. Mission equipment ó Check hoist, litters, radio and other equipment as required to support the mission. Refer to Chapter 4, MISSION EQUIPMENT, for equipment checks.

8-22. AREA 9 — FUSELAGE RIGHT SIDE.

e. Passenger seats and belts ó Check condition and security.

a. Fuselage ó Check condition. Check

f. Crewmember signal distribution panel (ICS) ó Check, set as desired.

Check condition, operation,

g. First aid kits ó Check condition and security.

b. Navigation and Position lights ó condition and security. c. Cabin doors ó and security.

d. Landing gear ó Check condition and security of cross tubes, skid, and skid shoe. Ground handling wheels removed.

8-23. AREA 10 — CABIN TOP. *a. Main rotor systems ó Check condition and security; check level of fluid in dampers, blade grips, and pillow blocks.

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Rev. 5

h. Fire extinguisher ó Check for charge (ensure seal is not broken) and security. *i. Crew and Passenger Briefing ó Complete as required. Refer to Passenger Briefing in Section VI.

8-25. BEFORE STARTING ENGINE. a. Seats and pedals ó Adjust.

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*b. Seat belts and shoulder harness — Fasten and tighten. c. Shoulder harness locks — Check operation and leave unlocked. *d. Control frictions and lock — Off. Set friction as desired. *e. Flight controls — Check for full travel. Center cyclic and pedals. Place collective full-down..

(4) Radio magnetic indicator — Set as desired; deviation card current. (5) Altimeters — Set to field elevation. (6) Airspeed indicators — Check indications. t. DC circuit breakers — IN, armament and special equipment.

except

for

*u. PITOT HTR switch — OFF. *v. POSITION light switches — As required STEADY for night; OFF for day.

f. AC circuit breakers — IN.

*w. ANTI COLL light switch — ON.

g. XMSN Chip — ON.

x. WIPERS switch — OFF.

h. Ignition key lock switch — ON.

y. CARGO REL switch — OFF. i. Avionics equipment — OFF; set on desired frequencies.

z. CABIN HEATING switch — OFF. aa.INST LTG switches — OFF.

j. GOV switch — AUTO.

ab.AC PHASE selector — AC.

k. ANTI-ICE switch — OFF.

ac. INVTR switch — OFF. l. FUEL switch — OFF. ad.MAIN GEN switch — ON and cover down. m. HYD CONT switch — ON.

ae. VM selector — ESS BUS.

n. FORCE TRIM switch — ON.

af. NON-ESS BUS switch — NORMAL ON.

o. Systems instruments — Check engine, transmission, and electrical system for static indications, slippage marks, and operating range limits.

ah.BAT switch — As required; ON for the battery start; OFF for GPU start.

p. COMPASS switch — As required. MAG for normal operation. Refer to Chapter 3 for free gyro mode operation.

*ai. Ground power unit — Connect for GPU start. The EXTERNAL POWER caution light should be illuminated.

ag.STARTER GEN switch — START.

aj. Free-air temperature gage — Check FAT and condition.

q. Radar Altimeter — OFF. r. Clock — Wound and running.

ak.FIRE warning indicator light — Test (15 seconds maximum).

s. Flight instruments — Check indications and set as follows:

al. BAGGAGE FIRE/DOOR TEST-TEST — This also will test ENG OIL FILTER indicator light.

(1) Turn and slip indicator — Check race full of fluid.

am.RPM warning light — Check illuminated.

(2) Magnetic compass — Full of fluid, deviation card current.

an.Master caution/warning lights — Check IFF Mode 4 and CODE HOLD, cargo release and marker beacon press to test lights.

(3) Vertical indications.

ao. Rotor brake — Check, light illuminated then off.

speed

indicators

—

Check

Rev. 5

8-7

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ap.Caution panel lights — Test and reset master caution light. aq.FUEL switch — ON. Check FUEL BOOST CAUTION light(s) out

*g. Gas producer (N1) — 68% to 72%. Hold a very slight pressure against the engine idle stop during the check. A slight raise in N1 may be anticipated after releasing pressure on throttle.

ar. Start fuel — ON. as. GOV RPM switch — DECR for 10 seconds. at. Throttle — Set for start. Check full travel and return to engine idle stop. Check operation of the engine idle stop, then close the throttle; position the throttle as near as possible (on decrease side) to the engine idle stop.

8-26. STARTING ENGINE.

The copilot attitude indicator should be caged and held momentarily as inverter power is applied. *h. Copilot attitude indicator — Cage and hold until an inverter is on. *i. INVTR switch — As required. SPARE for initial start. MAIN for thru-flight. The INVERTER caution light should be out. *j. Normal rising transmission oil pressures — Check.

In the event the engine has been motored with the throttle open and igniters not energized or there is any other reason to suspect the engine may be flooded with fuel, do the following: Prior to energizing igniters or attempting engine start, set the throttle full OFF then motor for 30 seconds to clear fuel from the engine and allow it to coast down to a full stop. Failure to clear fuel from the engine may result in engine damage.

*k. Ground power unit (GPU start) — Disconnect. The EXTERNAL POWER caution light should go out when the GPU access door is secured.

Check GPU disconnected prior to turning battery switch on. *l. BAT switch (GPU start) — ON.

*a. Fire guard — Posted.

8-27. ENGINE RUNUP.

*b. Rotor brake — Released. *c. Rotor blades — Check clear and untied. ★*d. Engine — Start as follows:

4

(1) Start switch — Press and hold; start clock for starter timing. Note DC voltmeter indication. A minimum of 24 volts should be indicated on the DC voltmeter before attempting start. Battery starts can be made when voltages less than 24 volts are indicated, provided the voltage does not drop below 14 volts with the starter energized. (2) Start fuel off at 750. (3) Release starter switch at 40% gas producer speed (N1) or after 35 seconds, whichever occurs first. Refer to Chapter 5 for starter limitations. e. Engine oil pressure — Check normal rising. *f. Throttle — Slowly advance past the engine idle stop to the engine idle position. Manually check the engine idle stop by attempting to close the throttle.

8-8

Rev. 7

For extended ground running (exceeding two minutes) N2 speed should be maintained at a minimum of 5500 rpm.

Full movement of cyclic at low rotor rpm may damage the main drive shaft. a. FORCE TRIM — Check as follows: (1) Press force trim button on the cyclic to ensure proper function. (2) Release button and place the FORCE TRIM switch to the OFF position. (3) Check cyclic and pedals for freedom of movement and tip-path for correlation with cyclic movement.

BHT PUB-92-004-10

b. Hydraulic system — check as follows: (1) Place the HYD CONT switch to the OFF position. Wait momentarily prior to removing hand from hydraulic control switch in case of hydraulic hardover. The HYD PRESSURE caution light should illuminate.

j. PITOT HTR — Check. Place the PITOT HTR switch in the ON position. Note loadmeter increase — then OFF. k. AC voltmeter — Check each phase for 115 ± 3 volts, leave in BC.

(2) Check controls for freedom of movement; ensure that the collective is full-down.

l. INVTR switch — OFF. Check for INST INVERTER caution light illumination and AC voltmeter indicates "zero". Switch to MAIN ON check caution light OFF.

(3) Place the HYD CONT switch to the ON position. The HYD PRESSURE caution light should be out.

m. AC voltmeter — Check each phase for 115 ± 3 volts, leave in AC.

(4) Place the FORCE TRIM switch to the ON position. c. Fuel boost/ Bleed Air and De-Ice. (1) Fuel Boost Pumps — Pull the RIGHT FUEL BOOST circuit breaker. The RIGHT FUEL BOOST caution light should illuminate and fuel pressure should remain in the normal range. RESET MASTER CAUTION. Pull the LEFT FUEL BOOST circuit breaker; the LEFT FUEL BOOST caution light should illuminate and pressure should be 0. Reset MASTER CAUTION. Press RIGHT FUEL BOOST circuit breaker in. The RIGHT FUEL BOOST caution light should be out and the fuel pressure should increase to the normal range. All caution lights should be out.

n. VM selector switch — Check all positions for DC voltage indication; leave in NON-ESS BUS position. Main generator voltage will be dependent upon the average ambient temperatures: 27 volts 28 volts 28.5 volts

26°C (80°F) and above 0°C to 26°C (32°F to 80°F) 0°C (32°F)

All standby generator positions should indicate 1.0 volt lower than the main generator voltage. *o. STARTER GEN switch — STBY-GEN.MAIN

d. BLEED AIR switch — ON; check EGT increase, OFF; EGT decrease.

p. MAIN GEN switch — OFF. The main generator loadmeter should indicate "zero" and standby generator loadmeter should indicate a load. Check for DC GENERATOR caution light illumination. DC voltage should indicate zero.

e. DE-ICE switch — ON; check EGT increase, OFF; EGT decrease.

q. NON-ESS BUS switch — MANUAL ON. Check DC voltmeter reading.

*f. Throttle — Slowly increase to full open. Engine rpm (N2) should stabilize at 6000 (6600 rpm for thru-flight) plus or minus 50 rpm. Throttle friction as desired.

r. VM selector switch — Check remaining positions for DC voltage indications; leave in MAIN GEN position.

*g. Engine and transmission instruments — Check readings are within operational limits. *h. Avionics — On as desired. i. Fuel quantity — Check by pressing the FUEL GAGE TEST switch until fuel quantity gage drops approximately 200 pounds, then release and check that gage returns to the original indication.

s. MAIN GEN switch — ON and guard closed. Check DC voltmeter. The DC GENERATOR caution light should be out. The main generator Loadmeter should indicate a load and the standby generator loadmeter should indicate "zero". *t. LOW RPM AUDIO switch — AUDIO. u. GOV INC-DEC switch — Slowly increase engine rpm (N2) to maximum, reading should be

Rev. 7

8-9

BHT PUB-92-004-10

6700, plus or minus 50 rpm. Set rpm at 6600. During governor INC-DEC check, the low rpm audio and warning light should go off when speed increases above 6100 ± 100 engine rpm and 300 ± 5 rotor rpm. v. Avionics — Perform operational check as necessary and set as desired. w. Clock — Set. *x. Attitude indicators — Set. *y. Heading indicators — Check that the heading indicators are aligned and correspond with magnetic compass. The annunciator may indicate alignment even though the heading indicator is 180° out of phase. Set heading indicator as required. *z. FORCE TRIM switch — As desired. *aa.Doors — Secured.

HIT Checks while operating in adverse conditions (e.g., dust, desert, coastal beach area, dry riverbeds) may be deferred (maximum of 5 flight hours at the discretion of the pilot in command until a suitable location is reached. ab.Health Indicator Tests (HIT) check — Perform on first flight of the day. The HIT provides the aviator with a go-no-go check for engine condition prior to takeoff. Perform HIT check as follows: (1) Maintain N2 at 6600 rpm. (2) Turn off all bleed air. (3) Turn helicopter into the wind and read free air temperature on cockpit FAT indicator. (4) Utilizing the HIT EGT log with baseline EGT data entered, locate the FAT in the first column nearest the free air temperature read on the cockpit FAT indicator. (5) Set N1% at the value indicated opposite this FAT. Allow EGT to stabilize. (6) Read EGT from indicator. (7) Compare this EGT with the baseline EGT adjacent to FAT and N1% utilized. (8) Record helicopter hours and different (±) between indicated EGT and baseline EGT. (9) If different between indicated EGT and baseline EGT is: (a) 20°C to 29°C, make an entry on DA Form 2408-13. (b) 30°C or greater/ make an entry on DA Form 2408-13; and do not fly the helicopter.

Rev. 6

Immediately prior to takeoff, the following checks shall be accomplished: *a. Rpm — 6600. *b. Systems — Check engine, transmission, electrical, fuel, systems indications are within operational limits. *c. Avionics — As required. *d. Crew, passengers, and mission equipment — Check. *e. Cargo hook — Check operation, if use is anticipated, perform the following steps: (1) Cargo — Secured; sling attached to cargo.

NOTE

8-10

8-28. BEFORE TAKEOFF.

(2) Ground required.

personnel

—

Positioned

as

(3) CARGO REL switch — Arm; check CARGO RELEASE ARMED light illuminates.

8-29. TAKEOFF TO HOVER. With the cyclic in a neutral position, increase collective with a smooth, positive pressure until the desired hovering altitude is reached. Apply pedal pressures to maintain heading as collective is increased. As the helicopter leaves the ground, make minor corrections with cyclic to ensure a vertical ascent and apply pedal pressures as necessary to maintain directional control. As the desired hover height is reached, adjust the flight controls as necessary to stabilize the helicopter at that height.

8-30. HOVERING TURNS. Apply pressure on the desired pedal to begin the turn, using pressure and counterpressure on pedals as necessary to maintain constant rate of turn. Coordinate cyclic to maintain the desired position over the selected point on the ground while maintaining altitude with collective.

8-31. SIDEWARD AND REARWARD HOVERING FLIGHT. From a stabilized hover, apply cyclic in the desired direction of flight to begin sideward or rearward movement. Maintain the desired heading with pedals and altitude with collective. Sideward and rearward airspeeds should be limited to 30 knots, which shall be estimated. To return to a stationary hover, apply cyclic opposite the direction of movement while coordinating collective and pedals to maintain the desired altitude and heading.

BHT PUB-92-004-10

8-32. HOVER/TAXI. From a stabilized hover, apply forward cyclic to begin forward movement of the helicopter. Maintain the desired heading with pedals and altitude with collective. Changes in direction should be made primarily with pedals to avoid excessive bank angles. To stop the forward movement, apply aft cyclic while coordinating collective and pedals to maintain the desired altitude and heading.

8-33. HOVER/TAXI CHECK. Perform the following checks at a hover/taxi. *a. Flight controls ó correct responses.

Check flight controls for

*b. Engine and transmission instruments ó Check. *c. Flight instruments ó Check as required. (1) VSI and altimeter ó descent.

WSPS lower cutter the forward c.g. high gross weight, high density altitude, transitional lift setting, and a tail wind increase the probability of ground contact. a. Cargo hook ó Check operation, if use is anticipated, perform the following steps: NOTE Avoid critical relative winds while performing external cargo operations. Refer to Critical Relative Wind Azimuths for Hover Flight illustration in Section in Section 7. b. helicopter at sufficient height to allow ground personnel to discharge static electricity and to attach cargo sling to cargo hook. NOTE

Indicate climb and

(2) Slip indicator ó Ball free in race. (3) Turn needle, heading indicator, and magnetic compass indicate turns left and right. (4) Attitude indicator ó Indicates nose high and low and banks left and right. (5) Airspeed indicator ó Check airspeed. *d. Power ó Check. The power will determine if sufficient power is available for takeoff and is performed by referring to the GO-NO-GO procedure in Figure 8-3, and by comparing torque required to hover with the predicted values from performance charts in Chapter 7.

8-34. LANDING FROM A HOVER. From a stabilized hover, decrease collective to begin a gradual descent to touchdown, making necessary corrections with pedals and cyclic to prevent movement over the ground. Upon contact with the ground, continue to decrease collective smoothly and steadily until the entire weight of the helicopter is on the ground. Apply cyclic as necessary to level rotor system.

8-35. TAKEOFF.

During take-off and at any time the helicopter skids are close to the ground, negative pitch attitudes (nose low) of 10∞or more can results in ground contact of the

Attachment of cargo sling to cargo hook can be observed by means of rear view mirror. c. Ascend vertically directly over cargo, then slowly lift cargo from surface. d. Check pedals for adequate directional control. e. Check torque required to hover with external load. f. Take off into wind if possible, allowing adequate sling load clearance over obstacles.

8-36. NORMAL. Align the helicopter with the desired takeoff course at a stabilized hover of approximately 3 feet (skid height), or an altitude permitting safe obstacle and terrain clearance. Smoothly apply forward cyclic to begin acceleration into effective translation lift. As the helicopter begins its forward movement, additional collective will be required to maintain altitude. Simultaneously adjust pedal pressure as necessary to maintain the desired heading. Control rate of acceleration and direction of flight with cyclic and altitude with collective. Continue to accelerate at the required altitude until effective translational lift has been attained, then begin a climb. Establish climb at the desired rate and airspeed. Continuous coordinated application of control pressures is necessary to maintain trim, heading, flight path, airspeed and rate of climb. Avoid downwind takeoffs when possible.

Rev. 6

8-11

BHT PUB-92-004-10

Figure 8-3. Go-no-go-takeoff data placard

8-12

BHT PUB-92-004-10

Refer to the height-velocity paragraph, Chapter 5, for avoid areas. The height-velocity paragraph assumes the availability of a suitable landing area in the event of engine failure. Since suitable landing areas are often not available, operating outside the avoid areas during takeoff and climb will provide t h e h i g h e s t m a r g i n o f s a f e t y. A d d i t i o n a l l y, autorotational landings can be made at minimum sink rate and low or zero forward ground speed, thereby minimizing damage and injuries, regardless of terrain. A normal takeoff may be made from the ground by aligning the helicopter with the desired takeoff

course on the ground and positioning the cyclic slightly forward of neutral. Smoothly increase co ll e ct iv e t o be g in a cl imb t o an a lt it u d e o f approximately three feet, (or an altitude permitting safe obstacle and terrain clearance) while s im ult a neous ly a ccel e ra ti ng t he heli copt er. Continue takeoff as above.

8-37. MAXIMUM PERFORMANCE. A takeoff that demands maximum performance from the helicopter is necessary because of various combinations of heavy helicopter loads, limited power and restricted

Rev. 6

8-12A/(8-12B blank)

BHT PUB-92-004-10

performance due to high density altitudes, barriers that must be cleared and other terrain features. The decision to use either of the following takeoff techniques must be based on an evaluation of the conditions and helicopter performance. The copilot (when available) can assist the pilot in maintaining proper rpm by calling out rpm and torque as power changes are made, thereby allowing the pilot more attention outside the cockpit.

cleared adjust helicopter attitude and collective as required to establish a climb at the desired rate and airspeed. Continuous coordinated application of control pressures is necessary to maintain trim, heading, flight path, airspeed, and rate of climb. Takeoff may be made from the ground by positioning the cyclic slightly forward of neutral prior to increasing collective.

8-38. COORDINATED CLIMB.

8-40. DELETED.

Align the helicopter with the desired takeoff course at a stabilized hover of approximately three feet (skid height). Apply forward cyclic smoothly and gradually while simultaneously increasing collective to begin a coordinated acceleration and climb. Adjust pedal pressure as necessary to maintain the desired heading. Maximum torque available should be applied (without exceeding helicopter limits) as the helicopter attitude is established that will permit safe obstacle clearance. The climbout is continued at the attitude and power setting until the obstacle is cleared. After the obstacle is cleared, adjust helicopter attitude and collective as required to established a climb as the desired rate and airspeed. Continuous coordinated application of control pressures is necessary to maintain trim, heading, flight path, airspeed, and rate of climb. Takeoff may be made from the ground by positioning the cyclic slightly forward of neutral prior to increasing collective.

8-41. COMPARISON OF TECHNIQUES.

8-39. LEVEL — ACCELERATION. Align the helicopter with the desired takeoff course at a stabilized hover of approximately three feet (skid height). Apply forward cyclic smoothly and gradually while simultaneously increasing collective to begin an acceleration at approximately 3 to 5 feet skid height. Adjust pedal pressure as necessary to maintain the desired heading. Maximum torque available should be applied (without exceeding helicopter limits) prior to accelerating through effective transitional lift. Additional forward cyclic will be necessary to allow for level acceleration to the desired climb airspeed. Approximately five knots prior to reaching the desired climb airspeed, gradually release forward cyclic and allow the helicopter to begin a constant airspeed climb to clear the obstacle. Care must be taken not to decrease airspeed during the climbout since this may result in the helicopter descending. After the obstacle is

Refer to Chapter 7, Performance Data, for a comparison of takeoff distances. Where the two techniques yield the same distance over a fifty-foot obstacle, the coordinated climb technique will give a shorter distance over lower obstacles and the travel acceleration technique will give a shorter distance over obstacles higher than fifty feet. The two techniques give approximately the same distance over a fifty-foot obstacle when the helicopter can just hover OGE. As hover capability is decreased the level acceleration technique gives increasing by shorter distance than the coordinated climb technique. In addition to the distance comparison, the main advantages of the level acceleration technique are:(1) It requires less or no time in the avoid area of the height velocity diagram; (2) performance is more repeatable since reference to attitude which changes with loading and airspeed is not required; (3) at the higher climbout airspeeds (30 knots or greater), reliable indicated airspeeds are available for accurate airspeed reference from the beginning of the climbout, therefore minimizing the possibility of descent. The main advantage of the coordinated climb technique is that the climb angle is established early in the takeoff and more distance and time are available to abort the takeoff if the obstacle cannot be cleared. Additionally large attitude changes are not required to establish climb airspeed.

8-42. SLINGLOAD. The slingload take-off requiring the maximum performance (when OGE hover is not possible) is similar to the level acceleration technique except the takeoff is begun and the acceleration made at 15 feet. Obstacle heights include the additional height necessary for a 15-foot slingload.

Rev. 5

8-13

BHT PUB-92-004-10 d. Landing light ó As required.

8-43. CROSSWIND TAKEOFF. A crosswind takeoff does not require a significantly different technique than a takeoff into the wind. The primary difference is the requirement to hold the lateral cyclic into the wind to prevent drift. For right crosswinds additional power is required because of more left pedal being applied to maintain the desired heading.

8-44. CLIMB. After takeoff, select the airspeed necessary to clear obstacles. When obstacles are cleared, adjust the airspeed as desired at or above the maximum rate of climb airspeed. Refer to Chapter 7 for recommended airspeeds.

8-45. CRUISE. When the desired cruise altitude is reached, adjust power as necessary to maintain the required airspeed. Refer to Chapter 7 for recommended airspeeds, power settings, and fuel flow.

8-46. DESCENT. Adjust power and attitude as necessary to attain and maintain the desired airspeed and rate during descent. Refer to Chapter 7 for power requirements at selected airspeeds and rates of descent. All checks of mission equipment that must be made in preparation for landing should be accomplished during descent. NOTE If cargo hook is used, then the following steps apply. (1) CARGO RELEASE switch ó approach.

ARM prior to

(2) Flight path and approach angle ó As required for wind and obstacle clearance. (3) Terminate approach to a high hover. When stabilized at a hover, descend slowly until cargo contacts surface. Maintain tension on sling. NOTE Release of cargo sling from cargo hook can be confirmed visually through rear view mirror. (4) CARGO RELEASE switch ó Press to release sling from cargo hook.

8-47. BEFORE LANDING. Prior to landing the following checks shall be accomplished: a. RPM ó 6600. b. Systems ó Check engine, transmission, electrical and fuel systems indications. c. Crew, passengers, and mission equipment ó Check.

8-14

Rev. 6

8-48. LANDING. 8-49. NORMAL APPROACH. The approach begins by adjusting power and attitude as required to establish an approach angle of approximately 8 to 10 degrees at an airspeed above the minimum rate of descent airspeed specified in Chapter 7. Maintain the entry airspeed until the apparent ground speed and rate of closure appear to increase. From this point, progressively decrease rate of descent and forward speed to stop all forward movement at an approximate three foot hover (or continue to the ground). As forward speed decreases below effective translational lift it will be necessary to gradually and smoothly increase collective to terminate the approach. Refer to paragraph 8-33, Landing from a Hover. Refer to Chapter 7 for airspeeds, rates of descent and power requirements. Refer to the Height Velocity paragraph, Chapter 5, for avoid areas during the approach.

8-50. STEEP APPROACH. A steep approach is used as necessary to clear obstacles in the approach path. It is executed in the same manner as the normal approach except for the slower airspeeds that must be maintained throughout the approach. Approach angles may vary from that of a normal approach to a vertical descent. Because of the increased power requirements during termination smooth and gradual collective movements are very important.

8-51. SHALLOW APPROACH. A shallow approach is used as necessary when instruments, tactical, or emergency requirements make it necessary to execute an approach at an angle less than that of a normal approach. It is executed in the same manner as a normal approach except that deceleration may be more rapid. Approach angles may vary from that of a normal approach to approximately zero.

8-52. RUNNING LANDING. A running landing is used during emergency conditions of hydraulic power failure and some flight control malfunctions and environmental conditions. The approach is shallow and flown at an airspeed that provides safe helicopter control. Airspeed is maintained as for a normal approach except that touchdown is made at an airspeed above effective translational lift. After ground contact is made, slowly decrease collective to minimize forward speed. If braking action is necessary, the collective may be lowered as required for quicker stopping.

BHT PUB-92-004-10

f. Throttle ó Off.

8-53. AFTER LANDING. a. Landing light ó As required.

g. Center pedestal switches ó OFF.

8-54. ENGINE SHUTDOWN.

h. FUEL switch ó OFF. i. Avionics ó OFF. j. Overhead switches ó OFF.

Deleted.

k. All electrical switches ó OFF except GEN, BAT, POSITION lights, and ANTI COLL light.

a. Collective ó Down. b. Throttle ó Reduce to engine idle. Allow EGT to stabilize for two minutes. Check N1 speed ó 68% to 72%. c. LOW RPM AUDIO switch ó OFF. d. FORCE TRIM switch ó ON.

l. Maintain rotor blades ó Tie down. m. Lights ó OFF. n. BAT switch ó OFF.

8-55. BEFORE LEAVING THE HELICOPTER.

e. STARTER GEN switch ó START.

a. Conduct a thorough walk-around inspection of the helicopter checking for damage, fluid leaks and levels. b. Mission equipment ó Secure.

If a rapid rise in EGT is noted, press the starter switch to motor the e n g i n e ( t h r o t t l e cl o s e d ) s t a b i l i z i n g temperature within limits.

c. Complete DA Forms 2408-12 and -13. d. Secure helicopter.

Section III. INSTRUMENT FLIGHT 8-56. INSTRUMENT FLIGHT — GENERAL. This helicopter is qualified for operation in instrument meteorological conditions. Flight handling qualities, stability characteristics, and range are the same during instrument flight as for visual flight. Navigation and communication equipment are adequate for instrument flight.

8-57. INSTRUMENT FLIGHT PROCEDURES. Refer to TM- 1-215, TM 1-225, TM 1-300, FLIP, AR 95-1, FAR Part 91, and procedures described in this manual.

a. Instrument Takeoff. Complete the normal checks prescribed in this chapter, to include a power check. An instrument takeoff should not be attempted with less than a hover out-of-ground effect capability. When ready for takeoff, align the helicopter with the takeoff course and set the RMI index to the takeoff heading. Set the attitude indicator miniature aircraft on the horizon. Using outside references to prevent movement of the helicopter, increase collective sufficiently to get the helicopter light on the skids. Then, while referring to the helicopter instruments, continue to increase collective smoothly and steadily until takeoff power is reached. Gradually adjust helicopter attitude to

Rev. 5

8-15

BHT PUB-92-004-10 place the attitude indicator miniature aircraft two bar-widths below the horizon, while maintaining a wings-level attitude. As the desired climb speed is reached, adjust helicopter attitude and power to maintain the desired airspeed and recommended rate of climb. b. Instrument Climb. Instrument climbs are normally performed at 80 KIAS; however, other airspeeds may be used when required. Refer to Chapter 7 for airspeeds, power required, rates of climb and fuel consumption. c. Instrument Cruise. There are no unusualflight characteristics during cruise in instrument meteorological conditions. d. Speed Envelope. Stability and handling characteristics are normal from maximum rate of climb airspeed throughout the entire speed envelope during instrument flight. Power settings during instrument flight should be in accordance with the cruise charts in Chapter 7. e. Communication and Navigation Equipment. Some special technique in the use of avionics are required for instrument flight. f. Instrument Descent. When a descent at slower than cruise speed is desired, slow the helicopter to the desired speed before initiating the descent. Normal descent or radar controlled descent to traffic pattern altitude can be made using cruise air speed. Normally, descent will be made at cruise speed by reducing power as required.

g. Holding. When holding enroute, it may be accomplished at cruise speed. Recommended speed when holding for an approach is 80 KIAS. For extended holding patterns at maximum endurance airspeeds, consult the appropriate cruise chart in Chapter 7. For descent in the holding pattern, decrease power and maintain the holding pattern airspeed. h. Instrument Approaches. Establish airspeed at 80 KIAS passing the initial approach fix. Continue with approach as depicted in instrument approach charts, or as instructed by the controller. Perform the before landing check prior to beginning final descent. The helicopter may be decelerated during the approach, but should not be flown at airspeeds below 60 KIAS while in instrument flight conditions.

8-58. NIGHT FLYING. Night flying is very closely related to instrument flying, and may often be conducted almost entirely under instrument conditions. Before takeoff, it is imperative to ensure that all lights, instruments, and avionics equipment are functioning properly. Generally, interior lighting should be kept to the minimum amount which will still allow complete visibility of all instruments and gages. Excessive cockpit lighting decreases outside visibility. Avoid using landing lights when in thick haze, smoke, or fog, as reflected light will reduce visibility and may affect depth perception. During ground operations, the helicopter should be hovered/taxied slowly, because it is difficult to judge actual ground speed and excess speeds may be developed without realizing it.

Section IV. FLIGHT CHARACTERISTICS 8-59. FLIGHT CHARACTERISTICS. 8-61. MAST BUMPING. 8-60. OPERATING CHARACTERISTICS. The flight characteristics of this helicopter in general are similar to other single rotor helicopters. NOTE Control movements should be made smoothly and kept to a minimum to prevent oscillation of sling load. CARGO REL switch ó OFF. Airspeed ó Within limits for adequate controllability of rotorcraft-load configuration. Fligh path ó As required to avoid flight with external load over any person, vehicle or structure.

8-16

Rev. 6

Abrupt inputs of flight controls cause excessive main rotor flapping, which may result in mast bumping and must be avoided. This condition occurs when the main rotor static stops contact the mast. It is most likely to occur when conducting slope operations and on rotor coast down in high wind conditions (natural or induced by other aircraft). It may be encountered in flight only if the helicopter flight envelope is exceeded.

8-62. COLLECTIVE BOUNCE. Collective bounce is a pilot-induced vertical oscillation of the collective system when an absolute friction (either pilot applied or control rigged) is less than seven pounds.

BHT PUB-92-004-10

It may be encountered in any flight condition by a rapid buildup of vertical bounce at approximately three cycles per second. The severity of the oscillation is such that effective control of the helicopter may become difficult to maintain. The pilot should apply and maintain adequate collective friction in all flight conditions. Should collective bounce be encountered relax pressure on the collective. (Do not "stiff arm" the collective). Make a significant pitch application either up or down, and increase collective friction.

8-63. BLADE STALL. a. Blade stall is caused by a high angle of attack on the retreating blade. (1) Blade stall is the result of numerous contributing factors such as gross weight, rotor rpm, airspeed, acceleration, and altitude. (2) The condition is most likely to occur athigher speeds and lower operating rpm; it alsofollows that the condition will occur sooner with higher values of altitude and gross weight. (3) One of the more important features of the two-bladed semi-rigid rotor system is its warning to the pilot of impending blade stall. (4) Prior to progressing fully into the stall region, the pilot will feel a marked increase in airframe vibration. Consequently, corrective action can be taken before the stall condition becomes severe. NOTE When rotor stall progresses into a severe state, feedback may occur, primarily in the cyclic. b. Blade stall ó Corrective Action. When blade stall is evident, the condition may be eliminated by accomplishing one or a combination of the following corrective: (1) Reduce collective.

8-64. MANEUVERING FLIGHT. Action and response of the controls during maneuvering flight are normal at all times when the helicopter is operated within the limitations set forth in this manual.

8-65. ROTOR CAPABILITIES. a. The helicopter is capable of delivering a maximum thrust commensurate with rotor-engine limitations and the density of the atmosphere in which it is operating. Maximum thrust can be utilized to obtain maximum airspeed, optimum rate of climb, or, at some reduce airspeed, the maximum maneuver potential. The pilot may employ the capabilities of the helicopter within maximum limitations and in accordance with the environment under which he is operating. b. A descending turn or autorotational turn at a given angle of bank and stabilized rate of descent imposes the same "G" load on the rotor. Hence, if the turn is too abrupt (tight) and rotor limits are exceeded, further application of controls will not check the rate of descent if the turn is continued. In order to alleviate this condition the pilot must roll out of the turn to reduce the rotor load and provide control response, and reduce rate of descent. c. The permissible bank angles vs altitudes and gross weights will affect the turning radius of the helicopter. A light gross weight helicopter turns within an area comparable in size to that contained within the boundaries of a football field. d. The same helicopter at normal gross weight and at a density altitude of 12,000 feet will require a much larger area to accomplish the same turn.

8-66. HOVERING CAPABILITIES. Refer to Chapter 7.

8-67. FLIGHT WITH EXTERNAL LOADS. The airspeed with external loads is limited by controllability.

(2) Reduce airspeed. (3) Increase operating rpm. (4) Descent to lower altitude. (5) Decrease the severity of the maneuver.

NOTE Exercise care, when carrying external loads, as the handling characteristics may be affected due to the size, weight, and shape of the external load.

Rev. 6

8-17

BHT PUB-92-004-10

8-68. FLIGHT CONTROLS COORDINATION. The most efficient maneuvering of this helicopter is obtained by the coordinated movement of the controls; coordinated control movement is more important for helicopter operation than it is for fixed-wing aircraft.

8-69. TYPES OF VIBRATION. a. The source of vibrations of various frequencies are the rotating and moving components on the helicopter, other components vibrate in response to an existing vibration. b. Rotor vibrations felt during in-flight or ground operations are divided in general frequencies as follows: (1) Extreme low frequency ó per revolution (pylon rock). (2) Low revolution.

frequency

ó

One

Less than one

or

two

per

(3) Medium frequency ó Generally, four, five, or six per revolution. (4) High frequency ó Tail rotor frequency or higher. c. Most vibrations are always present at low magnitudes. The main problem is deciding when a vibration level has reached the point of being excessive. d. Extreme low, and most medium frequency vibrations are caused by the rotor or dynamic controls. Various malfunctions in stationary components can affect the absorption or damping of the existing vibrations and increase the overall level. e. A number of vibrations are present which are considered a normal characteristic. Two per revolution is the most prominent of these, with four or six per revolution the next most prominent. There is always a small amount of high-frequency vibration present that may be detectable. Experience is necessary to learn the normal vibration levels. Sometimes the mistake is made of concentrating on feeling one specific vibration and concluding that the level is higher than normal.

8-18

Rev. 1

8-70. EXTREME LOW FREQUENCY VIBRATIONS. a. Extreme low-frequency vibration is usually limited to pylon rock. Pylon rocking, two to three cycles per second, is inherent with the rotor, mast, and transmission system. To keep the vibration from reaching noticeable levels, transmission mount dampers are incorporated to absorb the rocking. Malfunctions in the damper system will allow rocking to start. b. A quick check of the damper system may be made while in a hover. Moving the cyclic fore and aft at about one cycle per second will start the pylon rocking. The length of time it takes for the rocking to die out after the motion of the cyclic is stopped is indicative of the quality of the dampening. c. An abnormal continuation of rock during the check or a continued presence of rock during normal flight is an indication that something is wrong with the transmission mounts or dampers. d. Pylon rock may be aggravated when maximum power is reached, depending on condition of transmission dampers, engine and flight conditions. As much as plus or minus 5 psi torque oscillation may be observed. This condition may be corrected by reducing power approximately 1 percent N1.

8-71. LOW-FREQUENCY VIBRATIONS. a. Low-frequency vibrations are caused by the rotor itself. One per revolution vibrations are of two basic types, vertical or lateral. A one per revolution (1/rev) vertical is caused simply by one blade developing more lift at a given point than the other blade. A lateral vibration is caused by a spanwise unbalance of the rotor due to a difference of weight between the blades. The minor differences will affect flight but are compensated for by adjustments of trim tabs and pitch link settings. b. Generally, verticals felt predominantly in low power descent at moderate airspeeds (60-70 knots) are caused by a basic difference in blade lift and can be corrected by rolling the grip slightly out of track. Vertical vibrations felt in forward flight, increasing as airspeed increases, are usually due to one blade developing more lift with increased speed than the other (a climbing blade). This condition is corrected by adjustment of the trim tabs.

BHT PUB-92-004-10

8-72. LOW-FREQUENCY VIBRATION — VERTICAL. a. Associated with one per revolution vertical vibration is the intermittent one per revolution vertical. Essentially, this is a vibration initiated by a gust causing a momentary increase of lift in one blade giving a one per revolution vibration. b. The momentary vibration is normal; but if picked up by the rotating collective controls and fed back to the rotor causing several cycles of one per revolution, it becomes undesirable. Sometimes during steep turns one blade will "pop" out of track and cause a hard one per revolution vertical.

8-74. MEDIUM-FREQUENCY VIBRATIONS. a. Medium-frequency vibrations at frequencies of four or six revolution are another inherent vibration associated with most rotors. An increase in the level of these vibrations is caused by a change in the capability of the fuselage to absorb vibration, or a loose airframe component, such as the skids, vibrating at the frequency. b. Changes in the fuselage vibration absorption can be caused by such things as fuel level, external stores, structural damage, structural repairs, internal loading, or gross weight.

c. The condition is usually caused by too much differential tab in the blades and can be corrected by rolling one blade at the grip and removing some of the tab, (as much as can be done without effecting the ride in normal flight).

c. Abnormal vibration levels of this range are nearly always caused by something loose; either a regular part of the aircraft or part of the cargo or external stores. The vibration is felt as a rattling in the helicopter structure. The most common cause is loose skids.

8-73. LOW-FREQUENCY VIBRATION — LATERAL.

8-75. HIGH-FREQUENCY VIBRATIONS.

a. Should a rotor, or rotor component, be out of balance, a one per revolution lateral vibration will be present. This vibration is usually felt as a vertical due to the rolling motion it imparts to the helicopter, causing the pilot seats to bounce up and down out of phase; that is, the pilot goes up while the copilot goes down. An unusually severe lateral vibration can be felt as a definite sideward motion as well as a vertical motion. b. Lateral vibration existing due to an unbalance in the rotor are of two types; spanwise and chordwise. c. Spanwide unbalance is caused simply by one blade and hub being heavier than the other (i.e., an unbalancing along the rotor span).

a. High-frequency vibrations can be caused by anything in the ship that rotates or vibrates at a speed equal to or greater than that of the tail rotor. b. The most common and obvious causes; tail rotor balance and track. Pilot experience can help greatly in troubleshooting the cause of a highfrequency vibration, as a pilot who has experienced a vibration can often recognize the cause the next time he feels the same vibration. c. A comparison between the feel of the helicopter without excessive vibration and the helicopter with the vibration is helpful in precluding erroneous conclusion.

8-76. LOW "G" LOAD. (Less than 1.0)

d. A chordwise unbalance means there is more weight toward the trailing edge of one blade than the other. Both types of unbalance can be caused by the hub as well as the blades. e. Lateral vibrations are usually felt in a hover and in descending moderate airspeed turns and tend to disappear in forward flight, although many times a lateral can manifest itself as a vertical in forward flight. An out-of-ground effect hover is usually the best place to feel a lateral vibration and reducing the RPM to 6000 will often make the lateral more prominent. f. Two per revolution vibrations are inherent with two-bladed rotor systems and a low level of vibration is present.

Intentional flight below +0.5G is prohibited. Abrupt inputs of flig ht cont ro ls cause excessive main rotor flapping, which may result in mast bumping and must be avoided. Due to mission requirements, it may become necessary to lower the nose of the helicopter rapidly to recover from a pull up. Low "g" loads are placed on the helicopter with abrupt forward cyclic inputs. The

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8-19

BHT PUB-92-004-10 helicopter exhibits a tendency to roll to the right with forward cyclic input of sufficient force to create low "g" loads.

getting too large. the pilot must fly the helicopter into the air smoothly keeping excursions in pitch, roll and yaw low and not allowing any untrimmed moment.

NOTE When it becomes necessary to rapidly lower the nose of the helicopter, it is essential that the pilot monitor the changes in roll attitude as the cyclic is moved at a rate which will not produce excessive pitch and roll rates.

8-77. ROLLOVER CHARACTERISTICS. If a pilot finds himself in an accelerating right roll during a low "g" maneuver, he can recover within two to four seconds by pulling the cyclic sharply aft. The exact time to recover depends on how much bank the helicopter has developed before the correction is made. Aft cyclic produces a powerful left roll moment, and also increases rotor thrust and the pilot's roll control. Moving the collective or the pedals, will not speed recovery. Flight test have shown that adverse effects such as mast bumping and excessive tail rotor flapping may occur if these other controls are used. a. During normal or slope takeoffs and landings with some bank angle or side drift with one skid on the ground, the bank angle or side drift can cause the helicopter to get into the situation where it is pivoting about a skid. When this happens, lateral cyclic control response is more sluggish and less effective than for the free hovering helicopter. Consequently, if the bank angle (the angle between the helicopter and the horizon) is allowed to build up past about 15∞, the helicopter may enter a rolling maneuver that cannot be corrected with full cyclic and the helicopter may roll over on its side. In addition, as the roll rate and acceleration of the rolling motion increase, the angle at which recovery is still possible is significantly reduced. The critical rollover angle is also reduced for a right side down condition, crosswinds, lateral center of gravity offset and left pedal inputs. Refer to figure 8-4. b. When performing maneuvers with one skid on the ground, care must be taken to keep the helicopter trimmed, especially laterally. For example, if a slow takeoff is attempted and the tail rotor thrust contribution to rolling moment is not trimmed out with cyclic, the critical recovery angle may be exceeded. Control can be maintained if the pilot maintains trim, does not allow helicopter rates to become larger, and keeps the bank angle from

8-20

c. When performing normal takeoffs and landings on relatively level ground with one skid on the ground and thrust (lift) approximately equal to the weight, carefully maintain the helicopter position relative to the ground with the flight controls. Perform maneuvers smoothly and keep the helicopter trimmed so that no helicopter rates build up, especially roll rate. If the bank angle starts to increase to a larger angle (5∞ to 8∞) and full corrective cyclic does not reduce the angle, reduce collective to reduce the unstable rolling condition. d. When performing a slope takeoff and landing maneuver follow the published procedures, being careful to keep roll rates small. Slowly raise the downslope skid to bring the helicopter level and then lift off. (If landing, land on one skid and slowly lower the down-slope skid). If the helicopter rolls to the upslope side (5∞ to 8∞), reduce collective to correct the bank angle and return to level attitude and then start the takeoff procedure again. e. Collective is much more effective in controlling the rolling motion than lateral cyclic because it reduces the main rotor thrust (lift). A smooth, moderate collective reduction of less than approximately 40% (at a rate less than approximately full up to full down in two seconds) is adequate to stop the rolling motion with about two degrees bank angle overshoot from where down collective is applied. Care must be taken to not dump collective at too high a rate as to cause fuselage ó rotor blade contact. Additionally, if the helicopter is on a slope and the roll starts to the upslope side, reducing collective too fast creates a high rate in the opposite direction. When the up hill slope skid hits the ground, the dynamics of the motion can cause the helicopter to bounce off the upslope skid and the inertial can cause the helicopter to roll about down-slope skid and over on its side. Do not pull collective suddenly to get air borne as a large and abrupt rolling movement in the opposite direction will result. This movement may be uncontrollable.

BHT PUB-92-004-10

Figure 8-4. Forces acting on the helicopter

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If the helicopter reaches 15∞of bank with one skid on the ground and thrust (lift) approximately equal to the weight, the helicopter may roll over on its side.

When landing or taking off, with thrust (lift) approximately equal to the weight and one skid on the ground, keep the helicopter trimmed and do not allow helicopter rates to build up. Fly the helicopter smoothly off (or onto)the ground, carefullymaintaining trim.

Section V. ADVERSE ENVIRONMENTAL CONDITIONS 8-78. GENERAL. This section provides information relative to operation under adverse environmental conditions (snow, ice and rain, turbulent air, extreme cold and hot water, desert operations, mountainous and altitude operation) at maximum gross weight. Section II check list provides for operational requirements of this section.

8-79. COLD WEATHER OPERATIONS. Operation of the helicopter in cold weather or an arctic environment at temperatures as low as ñ40∞C (ñ40∞ F) presents no unusual problems if the operators are aware of those changes that do take place and conditions that may exist because of the lower temperatures and freezing moisture. At ambient temperatures lower than 40∞C (40∞F) the engine, drive system and hydraulic system must be preheated to ñ40∞ C (ñ40∞ F) or warmer prior to Engine Start. a. Inspection. The pilot must be more thorough in the walk-around inspection when temperatures have been at or below 0∞C (32∞F). Water and snow may have entered many parts during operations or in periods when the helicopter was parked unsheltered. This moisture often remains to form ice which will immobilize moving parts or damage structure by expansion and will occasionally foul electric circuitry. Protective covers afford majority protection against rain, freezing rain, sleet, and snow when installed on a dry helicopter prior to the precipitation. Since it is not practical to completely cover an unsheltered helicopter, those parts not protected by covers and those adjacent to cover overlap and joints require closer attention, especially after blowing snow or freezing rain. Remove accumulation of snow and ice prior to flight. Failure to do so can result in hazardous flight, due to aerodynamic and center of gravity disturbances as well as the introduction of snow, water, and ice into internal moving parts and electrical systems. The pilot should be particularly attentive to the main and tail rotor systems and their exposed control linkages.

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Rev. 5

If temperature is ñ44∞C (ñ47∞F) or below, the pilot must be particularly careful to monitor engine instruments for high oil pressure. b. Transmission. Check for proper operation by turning the main rotor opposite to the direction of rotation while observer watches the driveshaft to see there is no tendency for the transmission to wobble while the driveshaft is turning. If found frozen apply heat (do not use open flame, avoid overheating boot) to thaw the spherical couplings before attempting to start engine.

Prior to starting engine, on aircraft with improved partical separators and parked without covers installed, the upper half of the separator should be removed and inspected by maintenance personnel for ice and/ or snow. Any accumulation of these elements should be removed to prevent damage to engine. c. Checks. (1) Before exterior check 0∞C (32∞F) and lower. Perform check as specified in Section II, paragraphs 2-13 through 2-23. (2) Exterior check 0∞C (32∞F) to 54∞C (65∞F). Perform exterior check as outlined in Section II, paragraphs 2-13 through 2-23 plus the following checks.

Check that all surfaces and controls are free of ice and snow.

BHT PUB-92-004-10

NOTE Contraction of the fluids in the helicopter system at extreme low temperature causes indication of low levels. A check made just after the previous shutdown and carried forward to the walk around check is satisfactory if no leaks are in evidence. Filling when the system is coldsoaked will reveal an over-full condition immediately after flight, with the possibility of forced leaks at seals. (a) Main Rotor ó Check free of ice, frost, and snow. (b) Main Driveshaft ó Check for freedom of movement. (c) Engine air inlet ó Remove all loose snow that could be pulled into and block the engine intake during starting. (d) Oil cooling fan compartment ó Check oil cooler fan blades for ice. (3) Interior check ó all flights 0∞C (32∞F) to 54∞C (ñ65∞F). Perform check as specified in Section II, paragraph 2-24. (4) Interior check ó night flights 0∞C (32∞F) to ñ40∞C (ñ40∞F). External Power connected. Perform check as specified in Section II, paragraph 2-24. (5) Engine starting check 0∞C (32∞F) ñ54∞C (ñ65∞F). Determine that the compressor rotor turns freely. As the engine cools to an ambient temperature below 0∞C (32∞F) after engine shutdown condensed moisture may freeze engine seals. Ducting hot air from an external source through the air inlet housing will free a frozen rotor. Perform check as outlined in Section II, paragraph 2-26. NOTE During cold weather starting the engine oil pressure gage will indicate maximum (100 psi) and the transmission oil pressure will read above normal. The engine should be warmed up at engine idle until the engine oil pressure indication is below 100 psi and the transmission oil pressure is within operating limits. The time required for warmup is entirely dependent on the starting temperature of the engine and lubrication system.

(6) Hydraulic filter bypass indicator ó Reset if popped out. (7) Transmission OIL FILTER bypass indicator ó Reset if popped out. (8) Engine runup check. Perform the check as outlined in Section II, paragraph 2-27.

Control system checks should be performed with extreme caution when helicopter is parked on snow and ice. There is reduction in ground friction holding the helicopter stationary, controls are sensitive and response is immediate. e. Engine Starting Without External Power Supply. If a battery start must be attempted when the helicopter and battery have been cold-soaked at temperatures between ñ26∞C to ñ37∞C (ñ15∞F to -35∞F), preheat the engine and battery if equipment is available and time permits. Preheating will result in a faster starter cranking speed which tends to reduce the hot start hazard by assisting theengine to reacha self-sustaining speed (40% N1) in the least possible time. Electrical load may be reduced by leaving inverter lights and other electrical equipment OFF duringstart.

8-80. STOPPING THE ENGINE. At temperatures below minus 7∞C (20∞F), or if main rotor grip seal leakage was evident after previous engine shutdown, use the following procedure. a. Prior to engine shutdown, maintain 6000 to 6200 rpm, for one minute in minimum pitch. b. Make normal engine shutdown with collective in full down position. NOTE Do not use collective to decelerate rotor speed. c. At extreme low temperatures the time at minimum collective may have to be extended to allow seals to seat properly.

8-81. BEFORE LEAVING THE HELICOPTER. Open vents to permit free circulation of air. Install protective covers as required.

8-82. SNOW. a. Takeoff. Snow takeoff may be considered normal except for the following precautions that should be observed.

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BHT PUB-92-004-10

Under cold weather conditions, make sure all instruments have sufficient warm up time to ensure normal operation. Check for sluggish instruments before takeoff. (1) Select an area that is free of loose or powdery snow to minimize the restriction to visibility from blowing slow.

without extended hover in order to reduce the white-out condition that results from extended hovering over snow. This whiteout will usually occur on loose snow and can cause the pilot to lose all reference with the ground or any object he is approaching. If the object being used as a reference should become completely obscured, accomplish a go-around.

8-83. DESERT AND HOT WEATHER OPERATION. Problems encountered in desert operation are blowing dust/sand and high ambient temperature.

Due to air starvation, snow and ice accumulation during ground operation may be detrimental to the engine and hazardous to the helicopter and crew. Ground operation time should be minimized and FOD screen and particle separator must be inspected prior to takeoff. (2) Before attempting to take off make sure the landing gear skids are free and not frozen to the surface. (3) The first takeoff after a cold start should include a visual check of the ground surface for evidence of hydraulic leaks. This should be done under hovering power conditions. If hydraulic leaks are present, abort the mission. b. Landing ó Snow. Snow landing may be considered normal except for the following precautions that should be observed: (1) Select an area free of loose or powdery snow so that visibility will not be restricted by blowing snow. (2) Accomplish a normal landing to the ground. Limited visibility will result from swirling snow, when hovering is attempted before making a touchdown. (3) Anticipate loose powdery snow and crusts on all landings on snow. (4) Landings should always be made when visual ground reference can be maintained. The reference point should be kept forward and to the right so that it will be visible to the pilot at all times. NOTE When making an approach and landing on snow it should be one continuous operation

8-24

Rev. 5

a. Blowing dust and sand obscure vision. All takeoffs and landings should be made from or to the ground. b. High ambient temperature affect helicopter performance and engine and drive system component temperatures. Refer to Chapter 7.

8-84. TURBULENCE AND THUNDERSTORMS. Flight in turbulence and heavy rain which accompany thunderstorms should be avoided. If turbulence and thunderstorms are encountered inadvertently, use the following procedures:

8-85. TURBULENCE. a. In turbulence check that all occupants are seated with seat belts and harnesses tightened. b. Helicopter controllability is the primary consideration; therefore if control becomes marginal exit the turbulence as soon as possible. c. To minimize the effects of the turbulence encountered in flight the helicopter should be flown at an airspeed corresponding to maximum endurance airspeed. There will be a corresponding increase in control movements at the reduced airspeed. d. Reduce airspeed as much as practical to maintain safe flight but keep power on and maintain normal rotor rpm. e. Check that safety belts and harnesses are tightened. f. HTR switch ó ON. g. Power ó Adjust to maintain a penetration speed of 80 KIAS. h. Radios ó Turn volume down on any radio equipment badly affected by static. i. At night ó Turn interior lights to full bright to minimize blinding effect of lightning.

BHT PUB-92-004-10

j. Maintain a level attitude and constant power setting. Airspeed fluctuations should be expected and disregarded. k. Maintain the original heading, turning only when necessary. l. The altimeter is unreliable due to differential barometric pressures within the storm. An indicated gain or loss of several hundred feet is not uncommon and should be allowed for in determining minimum safe altitude.

8-86. LIGHTNING STRIKES. a. Although the possibility of a lightning strike is remote, with increasing use of all-weather capabilities the helicopter could inadvertently be exposed to lightning damage. Therefore static tests have been conducted to determine lightning strike effects on rotors. b. Simulated lightning tests indicate that lightning strikes may damage helicopter rotors. The degree of damage will depend on the magnitude of the charge and the point of contact. Catastrophic structural failure is not anticipated. However, lightning damage to hub bearings, blade aft section, trim tabs, and blade tips was demonstrated. Also adhesive bond separations occurred between the blade spar and aft section between the spar and leading edge abrasion strip. Some portions of blade aft sections deformed to the extent that partial or complete separation of the damaged section could be expected. Such damage can aerodynamically produce severe structural vibration and serious control problems which, if prolonged, could endanger the helicopter and crew.

NOTE Abnormal operating noises almost always accompany rotor damage, but loudness or pitch are not valid indications of the degree of damage sustained. d. If lightning strike occurs but there are no indications of damage to the helicopter, the following precautions are recommended to minimize further risk. (1) Reduce airspeed as much as practical to maintain safe flight but keep power on and maintain normal rotor RPM. (2) Proceed to the nearest suitable landing site, and descend with partial power, avoiding abrupt control inputs. (3) Do not autorotate, but accomplish precautionary landing, shutdown, and visually inspect rotors for damage before proceeding. (4) Record suspected lightning strike in DA Form 2408-13. e. If minor lightning damage is suspected but vibration indication is slight and no control problems appear, flight may be continued to a safe landing site, but avoid unnecessary delay in landing to assess damage. f. If lightning damage is moderately serious immediate landing is recommended. NOTE If mission requirements dictate resuming flight with damaged rotor blade, an aerodynamically smooth profile will minimize vibration and control problems.

Avoid flight in or near thunderstorms especially in areas of observed or anticipated lightning discharges. c. If lightning damage occurs, as indicated by control problems or vibration changes, especially abnormal noise, the pilot's assessment of the extent of damage, the mission requirements, and the demands of the current flight situation will determine the required action.

g. In the event severe lightning damage makes the helicopter difficult or impossible to control make an emergency landing or bail out.

8-87. ICE AND RAIN. NOTE Make entry of time flown in rain or around salt water in DA Form 2408-13.

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BHT PUB-92-004-10

a. In heavy rain, a properly adjusted wiper can be expected to clear the windshield adequately throughout the entire speed range. However, when poor visibility is encountered while cruising in rain, it is recommended that the pilot fly by reference to the flight instruments and the copilot attempt to maintain visual reference. Rain has no noticeable effect on handling or performance of the helicopter. NOTE If the windshield wiper does not start in LOW or MED position, turn the control to HIGH. After the wiper starts, the control may be set at the desired position. b. Intentional flight in known icing condition is prohibited. If icing conditions are encountered during flight every effort should be made to vacate the icing environment.

When operating at outside air temperatures of 4.4°C (40°F) or below, icing of the engine air inlet screens can be ex p ecte d . Ice accumulatio n o f inlet screens can be detected by illumination of the engine inlet air caution light. Continued accumulation of ice will result in partial or complete power loss. It should be noted that illumination of the engine inlet air caution light indicates blockage at the inlet screen only and does not reveal icing conditions in the sand and dust separator or on the FOD screen. Continuous flight in light icing conditions is not recommended because the ice shedding induces rotor blade vibrations, adding greatly to the pilot's work load. To preclude the possibility of icing, it is recommended that the right and left engine air inlet filters be removed from the cowling when it is anticipated that the helicopter will be flown under atmospheric conditions conducive to icing. (Do not remove the top filter.) . NOTE The use of engine de-ice on aircraft modified with the improved particle

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Rev. 7

separator (swirl tubes) should be limited to environmental condition in which OAT is 4° C (39°F) or below. c. If icing conditions become unavoidable the pilot should actuate the pitot heat, windshield defroster and the engine de-ice system. d. Flight tests in closely controlled icing conditions have indicated that the pilot can expect one or all of the following to occur. (1) Obscured forward field of view due to ice accumulation on the windscreens and chin bubbles. If the windshield defrosters fail to keep the windshield clear of ice, the side windows may be used for visual reference during landing. (2) One-per-rotor-revolution vibrations ranging from mild to severe caused by asymmetrical ice shedding from the main rotor system. The severity of the vibration will depend upon the temperatures and the amount of ice accumulation on the blades when the ice shed occurs. Flight test experience has shown that the possibility of an asymmetric ice shed occurring increases as the outside air temperature decreases. (3) An increase in torque required to maintain a constant airspeed and altitude due to ice accumulation on the rotor system.

If the ENGINE ICING light fails to illuminate in known icing conditions, or if for any other reason, the engine ice detector system is suspected to be inoperative, pull the ANTI-ICE ENG circuit breaker and check the ENGINE ICE DET light. Ensure DEICE switch is ON. If this light does not illuminate the pilot can be reasonably certain the engine ice detector system is inoperative. (4) Possible degradation of the ability to maintain autorotational rotor speed within operating limits. e. Severe vibrations may occur as a result of main rotor asymmetrical ice shedding. If icing conditions are encountered while in flight, land as soon as practical. All ice should be removed from the rotor system before attempting further flight.

BHT PUB-92-004-10

f. Control activity cannot be depended upon to remove ice from the main rotor system. Vigorous control movements should not be made in an attempt to reduce low-frequency vibrations caused by asymmetrical shedding of ice from the main rotor blades. These movements may induce a more asymmetrical shedding of ice, further aggravating helicopter vibration levels. g. If a 5 psi (or greater) torque pressure increase is required above the cruise torque setting used prior to entering icing conditions, it may not be

possible to maintain autorotational rotor speed within operational limits, should an engine failure occur. h. Ice shed from the rotor blades and/or other rotating components presents a hazard to personnel during landing and shutdown. Ground personnel should remain well clear of the helicopter during landing and shutdown and passengers, crewmembers should not exit the helicopter until the rotor has stopped turning.

Section VI. CREW DUTIES 8-87. CREW DUTIES.

(2) Altitude.

a. Responsibilities. The minimum crew to fly the helicopter is a pilot. Additional crewmembers as required may be added at the discretion of the commander. The manner in which each crewmember performs his related duties is the responsibility of the pilot in command.

(3) Time enroute. (4) Weather. d. Normal Procedures. (1) Entry and exit of helicopter

b. Pilot. The pilot in command is responsible for all aspects of mission planning preflight and operation of the helicopter. He will assign duties and functions to all other crewmembers as required. Prior to or during preflight the pilot will brief the crew on the mission performance data monitoring of instruments communications emergency procedures taxi and load operations.

(2) Seating.

c. Copilot. (when assigned). The copilot must be familiar with the pilots duties and the duties of other crew positions. The copilot will assist the pilot as directed.

(7) Smoking.

d. Crew Chief. The crew chief will perform all duties as assign by the pilot. e. Passenger Briefing. The following is a guide that should be used in accomplishing required passenger briefings. Items that do not pertain to a specific mission may be omitted. a. Crew Introduction. b. Equipment.

(3) Seat belts. (4) Movement in helicopter. (5) Internal communications. (6) Security of equipment.

(8) Oxygen. (9) Refueling. (10)Protective masks. (11)Parachutes. (12)Ear protection. (13)ALSE. (14)Danger areas -Refer to Figure 8-1. e. Emergency Procedures.

(1) Personal to include ID tags.

(1) Emergency exits.

(2) Professional.

(2) Emergency equipment.

(3) Survival

(3) Emergency landing/ditching procedures.

c. Flight Data. (1) Route

(4) Bail out.

8-88. DANGER AREAS. Refer To Figure 8-1.

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8-27/(8-28 blank)

BHT PUB-92-004-10

CHAPTER 9 EMERGENCY PROCEDURES Section I. HELICOPTER SYSTEMS 9-1.

HELICOPTER SYSTEMS.

This section describes the helicopter systems emergencies that may reasonably be expect to occur and pre sen ts th e pro ced ure s t o b e f ollo w ed . Emergency procedures are given in checklist form when applicable. A condensed version of these procedures is contained in the Operators and Crewmembers Checklist, BHT PUB 92-004-CL. Emergency procedures involving mission equipment are covered in Chapter 4, Mission equipment. Emergency operations of avionics equipment are covered when appropriate in Chapter 3, Avionics.

9-2. IMMEDIATE ACTION EMERGENCY PROCEDURES. Those steps that must be performed immediately in an emergency procedure are printed in bold type. These immediate action emergency procedures shall be committed to memory. NOTE The urgency of certain emergencies requires immediate and instinctive action by the pilot. The most important single consideration is helicopter control. All procedures are subordinate to this requirement.

9-3.

AFTER EMERGENCY ACTION.

After a malfunction of equipment has occurred, appropriate emergency actions have been taken and the helicopter is on the ground, and entry must be made in the Remarks Section of DA Form 2408-13 describing the malfunction.

9-4.

EMERGENCY EXITS.

Emergency exists are shown in figure 9-1. Emergency exist release handles are yellow and black striped. Activation of release handle will remove crew door or cargo door window.

9-5.

EMERGENCY ENTRANCE.

Crew/passenger removal is accomplished through the crew/cargo doors or through the windows with crash rescue equipment.

9-6.

EMERGENCY EQUIPMENT.

A fire extinguisher (figure 9-1) is mounted to the right of the pilot seat or the right center door post. 1. Extinguisher ó Remove from bracket. 2. Safety pin ó Pull. 3. Nozzle ó Aim at base of fire. 4. Handle ó Depress.

9-7.

ENGINE.

9-8.

FLIGHT CHARACTERISTICS.

The characteristics and reactions in this helicopter without power are similar to those of a normal power descent. Full control can be maintained with the helicopter in autorotational descent.

9-9.

ENGINE FAILURE AND AUTOROTATION.

The two conditions most likely to affect successful autorotational landings are the altitude and airspeed at which the helicopter is operating at the time of an engine failure. The main symptoms of either a partial power loss or complete engine failure are a sudden reduction in engine noise, a sudden drop of engine and rotor rpm, a left yaw resulting from the reduction in engine torque, and the total or partial lack of response to throttle movements. When a loss of engine power is detected, it is necessary to decrease the collective pitch and apply right pedal immediately in order to avoid a reduction in rotor rpm and to maintain a constant heading.

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9-1

BHT PUB-92-004-10

Figure 9-1. Emergency exits and equipment

9-2

BHT PUB-92-004-10

Under partial power conditions the engine may operate relatively smooth at reduced power or it may operate roughly and erratically with intermittent surges of power. In instances where a power loss is experienced without accompanying engine roughness or surging, the helicopter may sometimes be flown in a gradual descent at reduced power to a more favorable landing area. Under these conditions the pilot should always be prepared for a complete power failure and an immediate autorotational landing. In the event that a partial power condition is accompanied by engine roughness, erratic operational or power surging, take immediate action by closing the throttle completely and accomplish an autorotational landing.

9-10. MINIMUM RATE OF DESCENT — POWER OFF. Refer to figure 9-2.

9-11. MAXIMUM GLIDE DISTANCE — POWER OFF. Refer to figure 9-2.

9-12. ENGINE FAILURE. The minimum height for safe landing after engine failure chart depicts combinations of airspeed and height above the ground where a successful straight in autorotation may be made from level or hovering flight in the event of an engine failure. It is imperative that the techniques described in the subparagraphs be followed to achieve the capability shown by the height-velocity diagram. Delay in recognition of the failure, improper technique or excessive maneuvering to reach a suitable landing area reduces the probability of a safe touchdown. Flight conducted within the caution area of the height-velocity diagram exposes the helicopter to a high probability of damage despite the best efforts of the pilot. The following procedures must be observed when an engine failure occurs within the caution areas shown in Chapter 5, Height Velocity. 1. Area A. Forward cyclic should be applied in conjunction with lowering of collective to establish a nose-down pitch attitude of 25 degrees in order to reach stabilize on a descent airspeed as required in Chapter 7.

2. Area B. Forward cyclic should be applied as necessary in conjunction with lowering of collective to establish a nose-down pitch attitude of 20 degrees at 40 KIAS varying to no forward cyclic at 70 KIAS, in order to reach and stabilize on a descent airspeed as required in Chapter 7. NOTE Additional airspeed above the recommended dive speed increases rate of descent and should only be used as necessary to extend glide distance. 3. Area C. From conditions of low airspeed and low height, the deceleration capability is limited, and caution should be used to avoid striking the ground with the tail rotor. Initial collective reduction will vary from no reduction at zero airspeed and 15 feet to full down at 70 KIAS and 125 feet. Intermediate altitudes and airspeeds will require a partial reduction of collective in order to reach the termination prior to excessive rotor speed decay. Touchdown should be made in a slightly nosehigh attitude to reduce the ground run as much as possible. 4. When engine failure occurs, proceed as follows: a. Collective pitch ó Adjust as necessary; establish autorotational glide. (O)b. External appropriate. c.

stores/load

ó

Jettison

as

Land.

d. FUEL switch ó OFF. e. BAT switch ó OFF.

9-13. ENGINE DRIVESHAFT/CLUTCH FAILURE. A failure of the driveshaft/clutch will result in a complete loss of power to the rotor and a possible engine overspeed. The procedures for coping with this situation are the same as for an engine failure. If failure occurs proceed as follows: 1. Collective Pitch ó Adjust as necessary; establish autorotational glide.

9-3

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Figure 9-2. Glide distance chart

9-4

BHT PUB-92-004-10

2. Throttle — Close. 3. External appropriate.

stores/load

—

Jettison

as

4. Land. 5. FUEL switch — OFF. 6. BAT switch — OFF.

9-14. ENGINE GROUND OPERATIONS. 9-15. EMERGENCY START. An emergency start may be attempted if the engine fails to start normally and starting the engine is necessary because of an emergency situation. Normal before-starting checks must be completed prior to beginning the following procedure: 1. Throttle — Closed. 2. GOV switch — EMER.

engine failure cannot be determined in flight, the decision to attempt the start will depend on the altitude and time available, rate of descent, potential landing areas, and crew assistance available. Under ideal conditions approximately one minute is required to regain powered flight from time the attempted start is begun. If the decision is made to attempt an inflight start, proceed as follows: 1. Throttle — Closed. 2. STARTER GEN switch — START. 3. Fuel switchs — ON. 4. GOV switch — EMER. 5. Attempt start. a. Starter switch — Press. b. Throttle — Open slowly to 6600 rpm after N1 passes through 8 percent. Control rate of throttle application as necessary to prevent exceeding EGT limits.

3. Starter switch — Press. c. Starter switch — Release as N1 passes through 45 percent. After engine is started and powered flight is re-established continue flight with manual throttle control. Return the STARTER GEN switch to STBY GEN. Advance and reduce throttle slowly and monitor EGT closely when the GOV Switch is in the EMER position in order to avoid exceeding EGT limits or flameout. 4. Throttle — Open slowly to the engine idle position when N1 passes through 8 percent. 5. Starter switch — Release at 45 percent N1. 6. Throttle — Open slowly to 80 percent N1, then decrease slowly to engine idle.

6. Land as soon as possible.

9-17. ENGINE COMPRESSOR STALL. Engine compressor stall (surge) is characterized by a sharp rumble or a series of load sharp reports, severe engine vibration and a rapid rise in engine gas temperature (EGT) depending on the severity of the surge. Maneuvers requiring rapid or maximum power applications should be avoided. Should this occur the following steps should be accomplished.

7. GOV switch — Move to AUTO as N1 decreases from 80 percent to engine idle. Engine rpm may surge as the GOV switch is placed in the AUTO position; however, this is normal.

1. Reduce power.

9-16. ENGINE RESTART — DURING FLIGHT.

3. BLEED AIR switch — OFF.

After an engine failure in flight, resulting from a malfunction of the fuel control unit, an engine start may be attempted. Because the exact cause of

4. If stall continues, land as soon as practical.

2. DE-ICE switch — OFF.

5. After landing, accomplish a normal shutdown.

Rev. 7

9-5

BHT PUB-92-004-10

9-18. PARTIAL POWER. 1. Inlet Guide Vane Actuator Failure. — Closed or Open.

c. Throttle — Open slowly to normal operating RPM and continue flight with manual throttle control.

9-19. ENGINE OVERSPEED. a. Closed. If the guide vanes fail in the closed position a maximum of 20-25 psi of torque will be available. Although N1 may indicate normally, power applications above 20-25 psi will result in deterioration of N2 and rotor rpm while increasing N1. Placing the Governor Switch in the EMER position will not provide any increased power capability and increases the possibility of an N1 overspeed and an engine over-temperature. Should a failure occur, accomplish a shallow approach. b. Open. If the inlet guide vanes fail in the open position during normal flight it is likely that no indications will be evident. In this situation increased acceleration times will be experienced. As power applications are made from increasingly lower N1s, acceleration times will correspondingly increase. 2. Droop Compensator Failure. Droop compensator failure will be indicated when engine rpm fluctuates excessively during application of collective pitch. The engine will tend to overspeed as collective pitch is decreased and will underspeed as collective pitch is increased. In the event of failure of the droop compensator, proceed as follows: a. Collective pitch autorotational glide.

—

Down;

establish

b. Throttle — Engine idle. c.

1. Collective pitch — Increase to load the rotor and sustain engine rpm below the maximum operating limit. 2. Throttle — Reduce until normal operating rpm is attained. 3. Collective pitch — Down; establish autorotational glide when conditions permit. 4. Throttle — Engine idle. 5. GOV switch — EMER. 6. Throttle — Open slowly to normal operational rpm and continue flight with manual throttle control.

9-20. ENGINE UNDERSPEED. An engine underspeed is caused by a malfunctioning N2 governor. At low altitude/low airspeed the malfunction must be treated as an engine failure because of insufficient time and altitude to regain normal engine rpm with Emergency governor control. If an underspeed is experienced and altitude permits proceed as follows:

GOV switch — EMER.

d. Throttle — Open slowly to normal operating rpm and continue flight with manual throttle control. 3. Governor Control Failure. Failure of the GOV RPM switch or linear actuator results in the inability to control N2 RPM. Failure at normal operating RPM does not require action by the pilot. Should the failure occur below normal operating RPM an due to an emergency situation a takeoff must be made; proceed as follows: a. Throttle — Engine idle. b. GOV switch — EMER.

9-6

An engine overspeed is caused by a malfunctioning N2 governor. If an overspeed is experienced proceed as follows:

Rev. 5

1. Collective pitch autorotational glide.

—

Down;

establish

2. Throttle — Engine idle. 3. GOV switch — EMER. 4. Throttle — Open slowly to normal operating rpm and continue flight with manual throttle control. Because automatic acceleration, deceleration and overspeed controls are not provided with the GOV switch in the EMER position, control movements must be smooth to prevent overspeed, overtemperature or engine failure.

BHT PUB-92-004-10

9-21. ROTORS, TRANSMISSIONS, AND DRIVE SYSTEMS. 9-22. TAIL ROTOR MALFUNCTIONS. Because of the many different malfunctions that can occur it is not possible to provide a solution for every emergency. The success in coping with the eme rge ncy depe nds on quick an aly sis of the condition and selection of the proper emergency procedure. The following is a discussion of some typ es of malfu nction s, p robable eff ects, and corrective actions. a. Complete Loss of Tail Rotor Thrust. This is a situation involving a break in the drive system such as severed driveshaft, wherein the tail rotor stops turning or tail rotor controls fail with the tail rotor in a zero pitch condition, and no thrust whatsoever is delivered by the tail rotor. A failure of this type will always result in the nose of the helicopter turning to the right (left sideslip) and a left roll of the fuselage along the horizontal axis. It is likely that powered flight to a suitable area and execution of an autorotative approach is the prior emergency procedure. (1) In powered flight, the degree of sideslip and the degree of roll may be varied by changing airspeed and by varying power (throttle or pitch), but neither can be eliminated. Below an airspeed of approximately 30 to 40 knots, the sideslip angle may become uncontrollable, and the aircraft begins to revolve on its vertical axis. (2) In power-off flight (autorotation), the sideslip angle and the roll angle can be almost completely eliminated by maintaining an airspeed of 40 to 70 knots. When airspeed is decreased below approximately 20 to 40 knots, the fin stabilizing becomes negligible and the sideslip angle may become uncontrollable. Upon pitch application at touchdown, the fuselage will tend to turn in the same direction the main rotor is turning (left) due to an increase of friction in the transmission system. b. Fixed Pitch Setting. This is a malfunction involving a loss of control resulting in a fixed pitch setting, such as a severed control cable. Normally under these circumstances, the directional pitch setting that is in the tail rotor at the time the cable is severed will, to some degree, remain in the tail rotor

system. Whether the nose of the helicopter yaws left or right is dependent upon the amount of pedal (which is related to power) applied at the time the cable is severed. Regardless of pedal setting at the time of malfunction, a varying amount of tail rotor thrust will be delivered at all times during flight. (1) If the tail rotor pitch becomes fixed during an approach or other reduced power situation (right pedal applied), the nose of helicopter will turn right when power is applied, possibly to an even greater degree than would be experienced with complete loss of tail rotor thrust, and the overall situation may be even more hazardous. The best solution may not be to autorotate immediately. Whether a successful autorotation could be accomplished is not certain, and is dependent upon the amount of pitch applied at the time of malfunction. (2) If the tail rotor pitch becomes fixed during a takeoff or other increased power situation (left pedal applied), the nose of helicopter will turn left when power is reduced (as in leveling off with cruise power). This turn to the left upon power reduction will probably be to a greater degree than the left turn encountered in a lower powered situation. Under these circumstances, it appears that powered flight to an airfield and powered landing could be accomplished with little difficulty since the sideslip angle will probably be corrected when power is applied for touchdown. However, upon decreasing power to initiate the approach at destination the sideslip angle will increase and remain so increased during the approach, but should be corrected when touchdown power is applied. Due to sideslip increase upon reduction of power to initiate the approach, a higher than normal approach speed may be beneficial. In this instance, powered landing may be the best solution as it is unlikely that autorotation could be accomplished at all. (3) If the tail rotor pitch becomes fixed during normal cruise power settings, the helicopter reaction should not be so violent as in the previously described situations and, at speeds from 40 to 70 knots, the tail pylon should streamline with very little, if any, sideslip or roll angle. In this instance, autorotation may aggravate the situation because a reduction of power (torque) may then result in a right sideslip. It must be considered, however, that an increase in power at touchdown will result in a left sideslip if powered approach is used, although this sideslip should not be of a hazardous magnitude for touchdown.

Rev. 1

9-7

BHT PUB-92-004-10 c. Loss of Tail Rotor Components, Thereof. The gravity of this situation is dependent upon the amount of weight lost. If the loss is small, such as "aft of the 90 degree gearbox", the situation should be quite similar to "complete loss of tail rotor thrust." If more than that is lost, immediate autorotation may be the only solution of possible value. Any loss of this nature will result in a forward center of gravity shift, requiring aft cyclic control correction.

9-23. EMERGENCY PROCEDURES — IN FLIGHT. When an antitorque malfunction occurs, attempt to regain helicopter control in cruise flight. If the situation permits, a change in collective pitch (power) may be attempted as an aid in gaining maximum possible control (trim) of the helicopter under existing circumstances. An increase in collective pitch turns the nose right; decrease turns it left. Rolling off power (throttle) will turn the nose left. The courses of action available will normally be: a. An autorotational descent and landing should be accomplished where possible under most circumstances, except as described in paragraph (c) below. The autorotative technique to be used is described in paragraph (b) below. b. If a safe landing area is not immediately available, continue powered flight to a suitable landing area in powered flight with an airspeed dictated by the illuminations of the emergency condition. This airspeed should be that which is most comfortable to the pilot (between 40 and 70 knots) indicated. When the landing area is reached, make a full autorotative landing. During the descent, an airspeed above minimum rate of descent airspeed should be maintained and turns kept to an absolute minimum. If the landing area is a level paved surface, a run-on landing with a touchdown airspeed between 15 and 25 knots should be accomplished. If the field is unprepared, start to decelerate from about 75 feet altitude, so that forward groundspeed is at a minimum when the helicopter reaches 10 to 20 feet; execute the touchdown with a rapid collective pull just prior to touchdown in a level attitude with minimum ground roll (zero, if possible). c. If the tail rotor pitch is fixed in a "left pedal applied" position (tail rotor delivering thrust to the

9-8

Rev. 1

left) autorotative landing must not be attempted. The pilot should return to powered level flight at a comfortable airspeed which will be dictated by the degree of sideslip and roll; continue powered flight to the nearest improved landing area, and execute a run-on landing with power and a touchdown speed between 20 and 30 knots. Prior to final approach the throttle may be reduced to maintain engine rpm above 6000 manually (GOV switch ó AUTO). Upon decreasing power to initiate the approach to the landing area, the sideslip angle will increase for the duration of the approach, but should be corrected when touchdown power is applied.

9-24. EMERGENCY PROCEDURES — HOVER. a. If the tail rotor pitch is fixed in a "left pedal applied" position, gradually decrease collective pitch and land the helicopter. b. If total loss of tail rotor thrust is experienced, close the throttle immediately and accomplish an autorotational landing.

9-25. TRANSMISSION OIL — HOT OR LOW PRESSURE. If the transmission oil temperature XMSN OIL HOT caution light illuminates; limits on the transmission oil temperature gage are exceeded; XMSN OIL PRESS caution light illuminates; or limits on the transmission oil pressure gage are exceeded (low or high) accomplish an approach and landing with power immediately.

E ng i n e p o w e r m u s t b e m a i nt a i ne d throughout the approach and landing to aid in preventing seizure of gears in the transmission. Should transmission oil pressure drop to zero PSI, a valid cross reference can be made with the XMSN OIL PRESS warning light but not with the oil temperature indicators. The oil temperature gage and XMSN OIL HOT warning lights are dependent on fluid for valid indications.

BHT PUB-92-004-10

9-26. CHIP DETECTORS.

3. FUEL switches ó OFF.

9-30. FUSELAGE FIRE — GROUND.

Do not restart engine until cause of illumination has been corrected. If the CHIP DETECTOR caution light for the main transmission, engine, 42∞ and 90∞ gearboxes illuminates perform an approach and landing with power on as soon as possible.

9-27. FIRE. The safety of helicopter occupants is the primary consideration when a fire occurs; therefore, it is imperative that every effort be made by the flight crew to put the fire out. On the ground it is essential that the engine be shut down, crew and passengers evacuated and fire fighting begun immediately. If time permits, a "MAYDAY" radio call should be made before the electrical power is OFF to expedite assistance from airfield fire fighting equipment and personnel. If the helicopter is airborne when a fire occurs the most important single action that can be taken by the pilot is to land the helicopter.

If fire is observed in any part of the helicopter during ground operations proceed as follows: 1. Throttle ó Closed. 2. FUEL switch ó OFF. 3. BAT switch ó OFF. 4. Clear the helicopter of all passengers and crew immediately.

9-31. FUSELAGE FIRE — FLIGHT. If fire is observed in any part of the helicopter during ground operations proceed as follows: 1. Land immediately ó Perform a power-on approach and landing without delay. 2. Throttle ó Closed as soon as the helicopter is on the ground. 3. FUEL switch ó OFF. 4. BAT switch ó OFF.

Use fire extinguishers only in well ventilated area because the toxic fumes of the extinguishing agent may cause injury.

9-28. ENGINE FIRE. 9-29. HOT START — EMERGENCY SHUTDOWN. The following procedure is applicable during engine starting if flames are emitted from the tail pipe, EGT limits listed in Chapter 5 are exceeded, or it becomes apparent that it will be exceeded. 1. Starter switch ó Press. The starter switch must be held until EGT is in the normal operating range.

5. Clear the helicopter of all passengers and crew immediately.

9-32. ENGINE FIRE — FLIGHT. 9-33. LOW ALTITUDE. If the FIRE light illuminates or fire is observed in or around the engine compartment during flight at low altitude, proceed as follows: 1. Land immediately ó Perform a power-on approach and landing without delay. 2. Throttle ó Closed as soon as the helicopter is on the ground. 3. FUEL switch ó OFF. 4. BAT switch ó OFF.

2. Throttle ó Closed. The throttle must be closed immediately as the starter switch is depressed.

5. Clear the helicopter of all passengers and crew immediately.

Rev. 5

9-9

BHT PUB-92-004-10

9-34. CRUISE ALTITUDE. If the FIRE light illuminates or fire is observed in or around the engine compartment during flight at an altitude which will permit the execution of an autorotational descent and landing, proceed as follows:

9-37. ELECTRICAL FIRE — FLIGHT TO BE CONTINUED. If landing cannot be made immediately and flight must be continued, the defective circuits may be identified and isolated as follows: 1. Complete steps 1 through 3 above.

1. Collective pitch ó Down, autorotate. 2. Throttle ó Closed. 3. FUEL switch ó OFF. 4. Land ó Accomplish an autorotational descent and landing. 5. BAT switch ó OFF. 6. Clear the helicopter of all passengers and crew immediately.

9-35. ELECTRICAL FIRE. 9-36. ELECTRICAL FIRE — FLIGHT. In the event of electrical fire or suspected electrical fire in flight, proceed as follows:

2. Circuit breakers ó Out. As each of the following steps is accomplished, check for indications of the source of the fire. 3. MAIN GEN switch ó ON. 4. STARTER GEN switch ó STBY GEN. 5. BAT switch ó ON. 6. Circuit breakers ó In one at a time in the priority required. When malfunctioning circuit is identified, pull the applicable circuit breaker out.

9-38. ELECTRICAL FIRE — GROUND. In the event of electrical fire or suspected electrical fire during ground operations, proceed as follows: 1. Throttle ó Closed.

1. MAIN GEN switch ó OFF.

2. FUEL switch ó OFF.

2. STARTER GEN switch ó START.

3. MAIN GEN switch ó OFF.

3. BAT switch ó OFF.

4. STARTER GEN switch ó START. 5. BAT switch ó OFF.

4. Land immediately ó Perform a power-on approach and landing without delay. 5. Engine shutdown ó After landing complete as follows:

9-39. SMOKE AND FUME ELIMINATION.

a. BAT switch ó ON.

Smoke and/or toxic fumes entering the cockpit and cabin can be exhausted as follows:

b. Throttle ó Closed.

Doors, windows, and vents ó Open.

c.

FUEL switch ó OFF.

d. BAT switch ó OFF.

9-10

6. Clear the helicopter of passengers and crew immediately.

Rev. 5

Do not jettison doors in flight.

BHT PUB-92-004-10

9-40. FUEL SYSTEM.

2. Obtain personnel.

assistance

from

maintenance

9-41. FUEL BOOST PUMP FAILURE. 1. One Boost Pump. If the fuel pressure gage indicates a drop in pressure and/or one FUEL BOOST caution light illuminates, flight may be continued to a facility where the malfunction can be corrected. 2. Two Boost Pumps. If the fuel pressure gage indicates zero pressure and/or both FUEL BOOST caution lights illuminate, proceed as follows:

9-46A. CARGO ELECTRICALLY.

FAILS

TO

RELEASE

In the event cargo hook will not release cargo sling when CARGO RELEASE switch is pressed, proceed as follows: 1. Maintain tension on cargo sling. 2. Cargo release pedal — Push.

a. Descend to a pressure altitude of 4600 feet or less if possible. b. Land as soon as practical. No attempt should be made to troubleshoot the system while in flight. 3. Both fuel boost pumps must be operating before next flight.

9-42. FUEL FILTER CONTAMINATION.

9-47. MAIN GENERATOR MALFUNCTION. A malfunction of the main generator will be indicated by a zero indication of the Main Generator Loadmeter and DC GENERATOR caution light illumination. The standby generator will automatically pick up the load when the STARTER GEN switch is in the STBY GEN position. An attempt may be made to put the main generator back on line by accomplishing the following: 1. GEN & BUS RESET circuit breaker — IN.

If the FUEL FILTER caution light illuminates land as soon as practical. Correct problem before next flight.

9-43. ENGINE FUEL PUMP MALFUNCTION. If the ENGINE FUEL PUMP caution light illuminates accomplish a power-on approach and landing as soon as practical.

9-44. THROTTLE FAILURE — EMERGENCY SHUTDOWN. If the throttle cannot be closed, engine shutdown can be accomplished by placing the FUEL switch to the OFF position. The engine may continue to run from 1 1/2 to minutes after the switch is off.

9-45. ELECTRICAL SYSTEM.

2. MAIN GEN switch — RESET the ON. Do not hold the switch in the RESET position, if the main generator is not reclaimed or if it goes off the line again: 3. MAIN GEN switch — OFF, Continue the flight using the standby generator power.

9-48. HYDRAULIC. 9-49. HYDRAULIC POWER FAILURE. Hydraulic power failure will be evident when the force required for control movement increases; a moderate feedback in the controls is felt and/or the HYD PRESSURE caution light illuminates. Control mo v eme n t s w ill re s ul t in n orm al h el ico p te r response in every respect. In the event of hydraulic power failure:

9-46. ENGINE SHUTDOWN — ELECTRICAL FAILURE.

1. Airspeed — Adjust as necessary to attain the most comfortable level of control movements.

Normal engine shutdown cannot be accomplished after complete electrical failure. Therefore one of the following methods may be used:

2. HYD CONT circuit breaker — Out; if hydraulic power is not restored. 3. HYD CONT circuit breaker — In.

1. Connect another source of power (GPU or battery) if available.

4. HYD CONT switch — OFF.

Rev. 7

9-11

BHT PUB-92-004-10

condition could render the helicopter extremely hard to control. If the failure occurs proceed as follows: 1. On the Ground. Do not return the HYD CONT switch to the ON position for the remainder o f t h e f l i g h t . T hi s p r e v e n t s a n y possibility of a surge in hydraulic pressure and the resulting loss of control. 5. Land as soon as possible at an area that will permit a run-on landing with power. Maintain airspeed at or above effective translation lift until touchdown.

a. HYD CONT switch — ON. b. Shutdown procedure — Complete a normal engine shutdown. 2. In Flight. a. HYD CONT switch — ON. b. Land as soon as practical.

9-50. CONTROL STIFFNESS. Under certain conditions a failure within an irreversible valve within the boost system may cause extreme stiffness in the flight controls to the extent that controls are extremely hard to move. Should control stiffness occur proceed as follows: 1. HYD CONT switch — OFF the ON. Check for restoration of normal flight control movements. Repeat as necessary if control response is not restored. 2. HYD CONT switch — OFF if cycling the switch fails to restore controls to normal operation.

During simulated hydraulic failure operations the pilot turning the hydraulic control switch off will first alert the pilot at the controls. He should not remove his hand from the switch until he is sure the flight controls are functioning properly. Both pilots must have their hand on their respective cyclic as the switch is turned on or off.

9-52. LANDING AND DITCHING. 9-53. LANDING IN TREES.

Do not return the HYD CONT switch to the ON position for the remainder o f t h e f l i g h t . T hi s p r e v e n t s a n y possibility of a surge in hydraulic pressure and the resulting loss of control. 3. Land at an area that will permit a run-on landing with power. Maintain airspeed at or above effective translation lift until touchdown.

9-51. CYCLIC POWER. A cyclic hardover may occur when the HYD CONT switch is moved to the OFF position. This is the result of an irreversible valve remaining in an open position. Abrupt cyclic movement may be either left, right, or right rear depending on which irreversible fails. This

9-12

Rev. 5

A landing in trees should be made only when no other landing area is available. Select a landing area containing the least number of trees of minimum height. Decelerate to a zero ground speed at tree-top level and descend into the trees vertically applying collective pitch as necessary for minimum rate of descent. Prior to the main rotor blades entering the trees, ensure throttle is OFF and apply all of the remaining collective pitch.

9-54. DITCHING — POWER ON. If it becomes necessary to ditch the helicopter accomplish an approach to an approximate 3-foot hover above the water and proceed as follows: 1. Cockpit doors — jettison. 2. Cabin doors — open.

BHT PUB-92-004-10

3. Crew (except pilot) and passengers — exit.

9-56. FLIGHT CONTROL MALFUNCTIONS.

4. Hover a safe distance away from personnel.

Failure of components within the flight control system may be indicated through varying degrees of feedback, binding, resistance, or sloppiness. These conditions should not be mistaken for hydraulic power failure. In the event of a flight control malfunction, proceed as follows:

5. Throttle — close and execute an autorotation into the water. Apply full collective pitch to cushion landing. Apply rotor brake (if installed) to stop rotor. 6. Pilot — exit when the main rotor is stopped.

9-55. DITCHING — POWER OFF. If an engine failure occurs over water and ditching is imminent accomplish engine failure emergency procedures and proceed as follows: 1. Cockpit doors — jettison. 2. Cabin doors — open. 3. Land. Accomplish autorotational descent, decelerating to zero forward speed as the helicopter nears the water. Apply collective to cushion landing. Apply rotor brake to stop rotor. 4. Crew and passengers — Exit when the main rotor is stopped.

1. Land immediately — Perform a power-on approach and landing without delay. 2. Complete an engine shutdown immediately.

9-57. BAIL OUT. If a decision is made to bail out, proceed as follows: 1. Cockpit doors — Jettison. 2. Cabin doors — Open. 3. Trim helicopter and set force trim for a level attitude. 4. Bail out through nearest exit.

Rev. 7

9-13/(9-14 blank)

BHT PUB-92-004-10

INDEX Subject

Fig./Para.

Subject

Fig./Para.

A

B

Abbreviations, Definitions of ...............................7-10 AC Circuit Breaker Panel....................................2-61 AC Circuit Breaker Panel................................. F2-12 AC Power Supply System ....................................2-58 AC Power Indicators and Control .......................2-60 Acceleration ó Level ...........................................8-39 ADF Set AN/ARN-83 .............................................3-7 After Emergency Action.........................................9-3 After Landing .......................................................8-53 Air Induction System ...........................................2-17 Airspeed Calibration Chart .............................. F7-11 Airspeed Calibration ó Use of Chart .................7-37 Airspeed Indicators ..............................................2-72 Airspeed Limitations ...........................................5-11 Airspeed Operating Limits Chart ...................... F5-5 Allowable Loading................................................6-18 Allowable Loading............................................. F6-15 Alternate Troop Seat Placement ........................ F6-6 Altimeter...............................................................2-75 Altitude, Cruise ....................................................9-34 Altitude Indicators ...............................................2-76 Altimeter, Radar................................................ F3-10 Altimeter, Radar ó General................................3-10 AN/ARC-83 ADF Set..............................................3-7 AN/ARC-114A Control Panel.............................. F3-6 AN/ARC-114A FM Radio Set.................................3-6 AN/ARC-115, Control Panel ............................... F3-5 AN/ARC-115 VHF Radio Set .................................3-5 AN/ARC-159 UHF AM Control Panel ................ F3-4 AN/ARC-159 UHF Radio Set.................................3-4 Anticollision Light................................................2-63 Approach, Normal ................................................8-49 Approach, Shallow ...............................................8-51 Approach, Steep ...................................................8-50 Approved Commercial Fuels, Oils, and Fluids............................................................2-83 Area 1 ó Fuselage and Main Rotor ....................8-14 Area 2 ó Cabin Nose ...........................................8-15 Area 3 ó Fuselage Left Side ...............................8-16 Area 4 ó Aft Fuselage Left Side.........................8-17 Area 5 ó Tailboom Left Side...............................8-18 Area 6 ó Tailboom Aft.........................................8-19 Area 7 ó Tailboom Right Side ............................8-20 Area 8 ó Aft Fuselage Right Side ......................8-21 Area 9 ó Fuselage Right Side.............................8-22 Area 10 ó Cabin Top ...........................................8-23 Autorotation and Engine Failure ..........................9-9 Auxiliary Fuel System .........................................2-26

Bail Out ................................................................9-57 Basic Weight Checklist ó DD Form 365A............6-5 Battery ..................................................................2-54 Before Exterior Check..........................................8-12 Before Landing .....................................................8-47 Before Leaving the Helicopter .............................8-81 Before Starting Engine ........................................8-25 Before Takeoff.......................................................8-34 Blackout Curtains ................................................2-47 Blade Stall ............................................................8-63 Blanket, Receptacles Heated ...............................2-45 Bounce, Collective ................................................8-26 Briefings, Crew and Passenger..............................8-6 Bumping, Mast .....................................................8-61 C C-6533/ARC Signal Distribution Panel (ICS).............................................................3-3 C-6899/ARN-83 Direction Finder Control Panel ................................................................... F3-7 Cargo Center of Gravity Planning ......................6-13 Cargo Compartment............................................ F6-4 Cargo Compartment...............................................6-8 Cargo Hook .............................................................4-1 Cargo Loading ......................................................6-12 Cargo Loading Typical ........................................ F6-9 Cargo Suspension System, External .............................................................. F4-1 Cargo Tiedown Fitting Data ............................... F6-8 Caution Lights and Indicators.............................2-41 Caution Light, Hydraulic Pressure .....................2-36 Caution Panel .................................................... F2-13 Cautions, Warnings and Notes, Definition ...............................................................1-2 Center of Gravity Limitations ...............................5-9 Chapter 7 Index......................................................7-2 Checklist .................................................................8-9 Checklist Callout ..................................................8-11 Checks...................................................................8-10 Chip Detector........................................................9-26 Circuit Breaker Panel, AC ................................ F2-12 Circuit Breaker Panel, AC ...................................2-61 Circuit Breaker Panel, DC................................ F2-11 Circuit Breaker Panel, DC...................................2-57 Classification of Helicopter....................................6-2

Index 1

BHT PUB-92-004-10

INDEX (Cont) Subject

Fig./Para.

Climb ....................................................................8-44 Climb, Coordinated ..............................................8-38 Climb-Descent Chart .......................................... F7-8 Climb Performance (Maximum Torque ó 30 Minute Operation) Chart (Sheet 1 of 2) ............................................ F7-9 Climb Performance (Continuous Operation) Chart (Sheet 2 of 2) ......................... F7-9 Climb Performance Charts ..................................7-31 Cockpit Lights ......................................................2-67 Cold Weather Operations ....................................8-79 Collective Bounce .................................................8-62 Collective Control System ...................................2-29 Commercial Fuels, Oils and Fluids, Approved..............................................................2-83 Comparison of Techniques...................................8-41 Compartment Diagram....................................... F2-4 Compartment Diagram, Crew ...............................2-8 Compass Set, Gyromagnetic..................................3-8 Compass, Standby................................................2-78 Compressor Stall, Engine ....................................9-17 Conditions ó Climb Performance Charts ..........7-32 Console Panel Lights, Overhead .........................2-69 Control Panel AN/ARC-114A.............................. F3-6 Control Panel, AN/ARC-115 ............................... F3-5 Control Stiffeners.................................................9-50 Control Switch......................................................2-33 Controls and Indicators .......................................2-25 Controls and Instruments ...................................2-12 Cooling, Engine Compartment ............................2-16 Coordinated Climb ...............................................8-38 Covers, Ground Handling Equipment, Rotor Tiedowns, and Mooring Diagram.............2-85 Crew and Cargo Doors ...........................................2-9 Crew and Passenger Briefings ..............................8-6 Crew Compartment Diagram................................2-8 Crew Movement .................................................. F5-2 Crew Requirements, Minimum .............................5-4 Crosswind Takeoff................................................8-43 Cruise....................................................................8-45 Cruise Altitude .....................................................9-34 Cruise Chart, Clean Configuration, ñ30∞C, Seal Level to 2000 Feet (Sheet 1 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, ñ30∞C, 4000 Feet to 6000 Feet (Sheet 2 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, ñ30∞C, 8000 Feet to 14000 Feet (Sheet 3 of 13)..................................................... F7-7

Index 2

Subject

Fig./Para.

Cruise Chart, Clean Configuration, ñ15∞C, Seal Level to 6000 Feet (Sheet 4 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, ñ15∞C, 8000 Feet to 14000 Feet (Sheet 5 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, 0∞C, Seal Level to 6000 Feet (Sheet 6 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, 0∞C, 8000 Feet to 14000 Feet (Sheet 7 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, +30∞C, Seal Level to 6000 Feet (Sheet 10 of 13) ................................................... F7-7 Cruise Chart, Clean Configuration, +15∞C, 8000 Feet to 14000 Feet (Sheet 9 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, +15∞C, Seal Level to 6000 Feet (Sheet 8 of 13)..................................................... F7-7 Cruise Chart, Clean Configuration, +30∞C, 8000 Feet to 14000 Feet (Sheet 11 of 13) ................................................... F7-7 Cruise Chart, Clean Configuration, +45∞C, Seal Level to 6000 Feet (Sheet 12 of 13) ................................................... F7-7 Cruise Chart, Clean Configuration, +45∞C, 8000 Feet to 14000 Feet (Sheet 13 of 13) ................................................... F7-7 Cruise ó Conditions ............................................7-26 Cruise ó Use of Charts .......................................7-25 Curtains, Blackout ...............................................2-47 Cyclic Control System ..........................................2-28 Cyclic Hardover ....................................................9-51 D Danger Area ........................................................ F8-1 Data Basis ..............................................................7-6 Data Case .............................................................2-46 DC and AC Power Distribution ...........................2-51 DC Circuit Breaker Panel................................. F2-11 DC Circuit Breaker Panel....................................2-57

BHT PUB-92-004-10

INDEX (Cont) Subject

Fig./Para.

DC Power Indicators and Controls......................2-56 DC Power Supply System ....................................2-52 DD Form 365A ó Basic Weight Checklist............6-5 DD Form 365C..................................................... F6-2 DD Form 365C ó Basic Weight and Balance Record......................................................6-6 DD Form 365F..................................................... F6-3 DD Form 365F ó Weight and Balance Clearance Form F..................................................6-7 Definitions of Abbreviations ................................7-10 Defrosting and Heating System Controls ............................................................ F2-10 Descent .................................................................8-46 Desert and Hot Weather Operation ....................8-83 Detectors, Chip .....................................................9-26 Dimensions, Helicopter Principal....................... F2-2 Dimensions, Principal............................................2-3 Direction Finder Control Panel C6899/ARN-83.................................................... F3-7 Discrepancies, Performance...................................7-9 Ditching ó Power Off ..........................................9-55 Ditching ó Power On ..........................................9-54 Door Latch, Open ................................................ F5-4 Dome Lights .........................................................2-66 Doors, Crew and Cargo ..........................................2-9 Driveshafts ...........................................................2-40 Driveshaft/Clutch Failure, Engine ......................9-13 Droop Compensator .............................................2-22 E Electrical Circuit ..................................................2-37 Electrical Failure ó Engine Shutdown ..............9-46 Electrical Fire.......................................................9-35 Electrical Fire ó Flight .......................................9-36 Electrical Fire ó Flight to be Continued............9-37 Electrical Fire ó Ground ....................................9-38 Electrical System .................................................9-45 Electronic Equipment Configuration ................. F3-1 Electronic Equipment Configuration ....................3-2 Emergency Entrance..............................................9-5 Emergency Equipment ..........................................9-6 Emergency Exits ....................................................9-4 Emergency Exits and Equipment ...................... F9-1 Emergency Procedures ó Hover.........................9-24 Emergency Procedures ó In Flight ....................9-23 Emergency Start ..................................................9-15 Engine.................................................................. F2-8 Engine...................................................................2-15 Engine and Hydraulic Control Panels ............... F2-9

Subject

Fig./Para.

Engine Compartment Cooling .............................2-16 Engine Compressor Stall .....................................9-17 Engine Driveshaft/Clutch Failure .......................9-13 Engine Failure ......................................................9-12 Engine Failure and Autorotation ..........................9-9 Engine Fire ...........................................................9-28 Engine Fire ó Flight ...........................................9-32 Engine Fuel Control System................................2-18 Engine Fuel Pump Malfunction ..........................9-43 Engine Ground Operations..................................9-14 Engine Instruments and Indicators ....................2-23 Engine Limitations ................................................5-7 Engine Oil Supply System ...................................2-19 Engine Overspeed ................................................9-19 Engine Rating.........................................................5-8 Engine Restart ó During Flight.........................9-16 Engine Runup.......................................................8-27 Engine Shutdown .................................................8-54 Engine Shutdown ó Electrical Failure ..............9-46 Engine Starting ....................................................8-26 Engine Underspeed ..............................................9-20 Entrance, Emergency.............................................9-5 Environmental Restrictions.................................5-13 Equipment Configuration Electronic ....................3-2 Example Loading Problem ó External ........... F6-11 Example Loading Problem ó Internal ............ F6-10 Exceeding Operational Limits...............................5-3 Exits and Equipment, Emergency...................... F9-1 Exits, Emergency ...................................................9-4 Exterior Check (Figure 8-2) .................................8-13 Exterior Check Diagram ..................................... F8-2 External Cargo Rear View Mirror.......................2-48 External Cargo Suspension System ................... F4-1 External Loads, Flight with.................................8-67 External Power Receptacle ..................................2-53 Extreme Low Frequency Vibrations ...................8-70 F Failure, Engine .....................................................9-12 Filter, Hydraulic ...................................................2-35 Fire ........................................................................9-27 Fire Detector Warning System ............................2-79 Fire, Electrical ......................................................9-35 Fire, Electrical ó Flight ......................................9-36

Index 3

BHT PUB-92-004-10

INDEX (Cont) Subject

Fig./Para.

Fire, Electrical ó Flight to be Continued...........9-37 Fire, Electrical ó Ground ...................................9-38 Fire, Engine ..........................................................9-28 Fire, Engine ó Flight ..........................................9-32 Fire Extinguisher, Portable .................................2-13 Fire, Fuselage ó Flight .......................................9-31 Fire, Fuselage ó Ground.....................................9-30 First Aid Kits........................................................2-14 Flight Characteristics ..........................................8-59 Flight Characteristics ............................................9-8 Flight Controls Coordination ..............................8-68 Flight Control Malfunctions ................................9-56 Flight Plan..............................................................8-5 Flight With External Loads.................................8-67 FM Radio Set ó AN/ARC-114A ............................3-6 Forces Acting on the Helicopter ......................... F8-4 Force Trim System ...............................................2-31 Free-Air Temperature Indicator (FAT) ...............2-77 Fuel .......................................................................6-16 Fuel Boost Failure................................................9-41 Fuel Control System, Engine ..............................2-18 Fuel Data........................................................... F6-13 Fuel Filter Contamination...................................9-42 Fuels, Oils, and Fluids, Approved Commercial .........................................................2-83 Fuel Pump Malfunction, Engine .........................9-43 Fuel Supply System .............................................2-24 Fuel System..........................................................9-40 Fuel System, Auxiliary ........................................2-26 Fuels, Types and Uses .........................................2-84 Fuselage..................................................................2-5 Fuselage Fire ó Flight ........................................9-31 Fuselage Fire ó Ground .....................................9-30 G General Arrangement ......................................... F2-1 General Conditions ................................................7-8 Gearboxes .............................................................2-39 General Arrangement ............................................2-2 Generator Malfunction, Main..............................9-47 Glide Distance Chart .......................................... F9-2 Glide Distance, Maximum ó Power Off .............9-11 Go-No-Go Takeoff Data Placard ............................8-3 Governor RPM Switch .........................................2-21 Gross Weight for Safe Pedal Margin Chart ....... F5-3 Ground Handling Equipment, Covers, Rotor Tiedowns and Mooring Diagram ..................... F2-15 Ground Handling Equipment, Covers, Rotor Tiedowns, and Mooring Diagram .......................2-85

Index 4

Subject

Fig./Para.

Ground Operations, Engine .................................9-14 Gyromagnetic Compass Indicator (RMI) ........... F3-8 Gyromagnetic Compass Set ...................................3-8 H Hardover, Cyclic ...................................................9-51 Heated Blanket Receptacles ................................2-45 Heater, Pitot .........................................................2-44 Heating and Defrosting System ..........................2-50 Heating and Defrosting System Controls ........ F2-10 Height Velocity .....................................................5-14 Helicopter Principal Dimensions........................ F2-2 Helicopter Station Diagram................................ F6-1 Helicopter Station Diagram...................................6-3 Helicopter Systems ................................................9-1 High-Frequency Vibrations .................................8-75 Hot Start ó Emergency Shutdown.....................9-29 Hover Chart......................................................... F7-5 Hover ó Conditions .............................................7-20 Hover ó Use of Charts ........................................7-19 Hovering Capabilities ..........................................8-66 Hover Check .........................................................8-32 Hover ó Emergency Procedures.........................9-24 Hover/Taxi ............................................................8-31 Hovering Flight Sideward and Rearward...........8-30 Hovering Turns ....................................................8-29 Hydraulic ..............................................................9-48 Hydraulic & Engine Control Panel .................... F2-9 Hydraulic Filter....................................................2-35 Hydraulic Power Failure......................................9-49 Hydraulic Pressure Caution Light ......................2-36 I Ice and Rain..........................................................8-86 Idle Fuel Flow Chart......................................... F7-10 Idle Fuel Flow ó Conditions ...............................7-35 Idle Fuel Flow ó Use of Charts ..........................7-34 Ignition ó Starter System...................................2-20 Immediate Action Emergency Checks ..................9-2 Indicators and Caution Lights ............................2-41 Indicators and Controls .......................................2-25 Induction System, Air ..........................................2-17 Instrument Flight ó General .............................8-56 Instrument Flight Procedures .............................8-57

BHT PUB-92-004-10

INDEX (Cont) Subject

Fig./Para.

Instrument Lights................................................2-68 Instrument Markings ......................................... F5-1 Instrument Markings ............................................5-5 Instrument Panel ó UH-1HRS ......................... F2-6 Instruments and Controls ...................................2-12 Instruments and Indicators, Engine ...................2-23 Interior Check ó Cabin.......................................8-24 Inverters ...............................................................2-59 L Landing.................................................................8-48 Landing and Ditching ..........................................9-52 Landing From a Hover.........................................8-33 Landing Gear System ............................................2-7 Landing in Trees ..................................................9-53 Landing Light.......................................................2-64 Landing, Running ................................................8-52 Level ó Acceleration ...........................................8-39 Lightning Strikes .................................................8-85 Limitations, Airspeed...........................................5-11 Limitations, Center of Gravity ..............................5-9 Limitations, Engine ...............................................5-7 Limitations, Rotor ..................................................5-6 Limitations, Weight .............................................5-10 Limits......................................................................7-4 Limits Chart, Airspeed Operation ..................... F5-5 Litter Installation ó Typical ........................... F6-12 Litter Racks ..........................................................6-15 Loading, Allowable...............................................6-18 Loading, Allowable............................................ F6-15 Loading and Unloading of Other Than General Cargo............................................6-14 Loading and Unloading, Personnel .....................6-10 Loading, Cargo .....................................................6-12 Loading Charts.......................................................6-4 Loading Problem ó External, Example .......... F6-11 Loading Problem ó Internal, Example ........... F6-10 Low Altitude .........................................................9-33 Low-Frequency Vibrations ..................................8-71 Low-Frequency Vibration ó Lateral ..................8-73 Low-Frequency Vibration ó Vertical..................8-72 Low ìGî Load (Less than 1.0) ..............................8-76 M Main Standby Starter ó Generator ...................2-55 Main Generator Malfunction...............................9-47 Main Rotor............................................................2-42 Main Rotor Blade .................................................2-86 Malfunctions, Flight Control ...............................9-56

Subject

Fig./Para.

Malfunctions, Tail Rotor ......................................9-22 Maneuvering Flight .............................................8-64 Maneuvers, Prohibited.........................................5-12 Marker Beacon Controls ..................................... F3-9 Marker Beacon Receiver ........................................3-9 Mast Bumping ......................................................8-61 Master Caution System .......................................2-80 Maximum Glide Distance ó Power Off ..............9-11 Maximum Performance........................................8-37 Medium-Frequency Vibrations ............................8-74 Minimum Crew Requirements ..............................5-4 Minimum Height for Safe Landing After Engine Failure Chart ............................... F5-6 Minimum Rate of Descent ó Power Off .............9-10 Maximum Torque Available (30 Minute Operation) Chart ................................................ F7-3 Mission Planning....................................................8-1 Mirror, External Cargo Rear View ......................2-48 Moments, Personnel .............................................6-11 Movement, Crew ................................................. F5-2 N Navigation Lights.................................................2-62 Night Flying .........................................................8-58 Normal Approach .................................................8-49 Notes, Definition ó Warnings, Cautions and .........................................................1-2 O Oil..........................................................................6-17 Oils and Fluids, Approved Commercial Fuels ...............................................2-83 Oil Data ............................................................. F6-14 Oil Level Light, Transmission .............................2-71 Oil Supply System, Engine ..................................2-19 Open Door Latch ................................................. F5-4 Operational Limits, Exceeding ..............................5-3 Operating Characteristics ...................................8-60 Operating Limits and Restrictions........................8-2 Operating Procedures and Maneuvers .................8-7 Overhead Console Panel Lights ..........................2-69 Overspeed, Engine ...............................................9-19

Index 5

BHT PUB-92-004-10

INDEX (Cont) Subject

Fig./Para. P

Partial Power........................................................9-18 Pedestal Lights.....................................................2-70 Performance ...........................................................8-4 Performance Discrepancies ...................................7-9 Performance Planning Card ................................7-12 Performance Planning Card ............................... F7-1 Performance Planning Sequence.........................7-13 Personnel Loading and Unloading ......................6-10 Personnel Moment .............................................. F6-7 Personnel Moments..............................................6-11 Personnel Seats ....................................................2-11 Pilot/Copilot Seats............................................... F2-5 Pilot/Copilot Seats................................................2-10 Pilot Station Diagram......................................... F2-7 Pitot Heater..........................................................2-44 Portable Fire Extinguisher ..................................2-13 Power Application ................................................8-40 Power Available ó Chart Differences.................7-15 Power Available ó Conditions ............................7-17 Power Distribution, DC and AC ..........................2-51 Power Failure, Hydraulic.....................................9-49 Power Supply and Circuit Breakers................... F3-3 Power Supply System, AC ...................................2-58 Power Supply System, DC ...................................2-52 Principal Dimensions.............................................2-3 Prohibited Maneuvers .........................................5-12 R Radar Altimeter ................................................ F3-10 Radar Altimeter ó General ................................3-10 Radius, Turning .....................................................2-4 Rain and Ice .........................................................8-86 Rate of Descent, Minimum Power Off.................9-10 Receptacle, External Power .................................2-53 Reservoir and Sight Glass ...................................2-34 Restart, Engine ó During Flight........................9-16 Restrictions and Operating Limits .......................8-2 Rollover Characteristics ......................................8-77 Rotor Capabilities ................................................8-65 Rotor Brake, Main................................................2-86 Rotor Limitations ...................................................5-6 Rotor, Main ...........................................................2-42 Rotor, Tail .............................................................2-42 Rotor Tiedowns and Mooring Diagram, Ground Handling Equipment Covers ................2-85 Rotors, Transmission, and Drive Systems................................................................9-21

Index 6

Subject

Fig./Para.

RPM Governor Switch .........................................2-21 RPM High-Low Limit Warning System ..............2-81 Running Landing .................................................8-52 Runups, Engine ....................................................8-27 S Starter-Generator, Main and Standby ................2-55 Seat Placement, Alternate Troop ....................... F6-6 Seat Placement, Troop ........................................ F6-5 Seats, Personnel ...................................................2-11 Seats, Pilot/Copilot ...............................................2-10 Seats, Troop ............................................................6-9 Searchlight ...........................................................2-65 Servicing ...............................................................2-82 Servicing Diagram ............................................ F2-14 Shallow Approach ................................................8-51 Shutdown, Engine ................................................8-54 Shutdown, Hot Start ó Emergency....................9-29 Sideward and Rearward Hovering Flight....................................................................8-30 Sight Glass and Reservoir ...................................2-34 Signal Distribution Panel (ICS) C-6533/ARC ...........................................................3-3 Signal Distribution Panel (ICS) ......................... F3-2 Slingload ...............................................................8-42 Smoke and Fume Elimination.............................9-39 Snow......................................................................8-82 Specific Conditions .................................................7-7 Stall, Blade ...........................................................8-63 Standby Compass .................................................2-78 Station Diagram, Helicopter............................... F6-1 Station Diagram, Helicopter..................................6-3 Station Diagram, Pilot ........................................ F2-7 Start, Emergency .................................................9-15 Starter-Ignition System .......................................2-20 Starting Engine ....................................................8-26 Stopping the Engine.............................................8-80 Steep Approach ....................................................8-50 Stiffness, Control..................................................9-50 T Tailboom .................................................................2-6 Tail Rotor Malfunctions .......................................9-22 Tail Rotor ..............................................................2-43

BHT PUB-92-004-10

INDEX (Cont) Subject

Fig./Para.

Tail Rotor Control System ...................................2-30 Takeoff ..................................................................8-35 Takeoff Chart ...................................................... F7-6 Takeoff ó Conditions...........................................7-23 Takeoff, Crosswind ...............................................8-43 Takeoff to Hover ...................................................8-28 Takeoff ó Use of Charts......................................7-22 Taxi/Hover ............................................................8-31 Temperature Conversion Chart ....................... F7-17 Temperature Indicator (FAT) Free Air ...............2-77 Throttle Failure ó Emergency Shutdown..........9-44 Tiedown Fitting Data, Cargo.............................. F6-8 Torque Available (Continuous Operation ó ECS Off) Chart ................................................... F7-4 Transmission ........................................................2-38 Transmission Oil Level Light..............................2-71 Transmission Oil ó Hot or Low Pressure ..........9-25 Troop Seats.............................................................6-9 Troop Seat Placement ......................................... F6-5 Turbulence and Thunderstorms..........................8-84 Turn and Slip Indicators .....................................2-73 Turning Radius ................................................... F2-3 Turning Radius ......................................................2-4 Types and Uses of Fuels ......................................2-84 Types of Vibration ................................................8-69

Subject

Fig./Para. U

Underspeed, Engine .............................................9-20 UHF AM AN/ARC-159 Control Panel ................ F3-4 UHF Radio Set AN/ARC-159.................................3-4 Use of Charts ..........................................................7-5 Use of Climb Performance Charts.......................7-31 Use of Climb-Descent Chart ................................7-28 W Warnings, Cautions, and Notes, Definition ...............................................................1-2 Warning System, Fire Detector ...........................2-79 Weight and Balance Record, DD Form 365C .............................................................6-6 Weight, Balance, and Loading ...............................8-3 Weight Limitations...............................................5-10 X Y Z

Rev. 3

Index 7/(Index 8 blank)