Bht-412-Fm-4 Bell 412ep Flight Manual

BHT-412-FM-4

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O NOTICE

Please inser t attached BHT-412-FM-4, REVISION 24 dated 05 NOVEMBER 2002 into Flight Manual in accordance with Log of Pages.

POST OFFICE BOX 482

FORT WORTH, TEXAS 76101

23 JUNE 1994 REVISION 24 — 05 NOVEMBER 2002

BHT-412-FM-4

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G ROTORCRAFT O N I FLIGHT MANUAL R W O E I F V NLY O BHT 36087 AND SUBSEQUENT

POST OFFICE BOX 482


BHT-412-FM-4

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O ROTORCRAFT FLIGHT MANUAL

BHT 36087 AND SUBSEQUENT

THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS

COPYRIGHT NOTICE COPYRIGHT 2002 BELL ® HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON CANADA LTD. ALL RIGHTS RESERVED

POST OFFICE BOX 482


23 JUNE 1994 REVISION 24 — 05 NOVEMBER 2002

BHT-412-FM-4

NOTICE PAGE

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O PROPRIETARY RIGHTS NOTICE

Manufacturer’s Data portion of this supplement is proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron Inc.

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 76101-0482 NP

Rev. 24

05 NOV 2002

BHT-412-FM-4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODEL

ROTOR CRAFT FLIGHT MANUAL

(f

U

36087 AND SUBSEQUENT

GHIORARY REVISION MAST TFRQf SIGNAL CONDITIONER M 32-97-90 -v/ This Temporary Ren perfrmance and pro compliance with ASB 412

current published ace data after

ew

?

Attach these temporary revisê pages in light Manual after con

like numbered 412-97-90.

DO NOT remove existing pages rua

Infornaton contained herein suppenn Flight Manual. For Umitations, Proced Data not contained in this supplement, Manual.

basic

ac

THIS MANUAL SHALL BE IN THE HELICOPTER DURING

Bell HelicoptertmE lr'l' l

COfRIIT NOICE COpYRulIr 2000 BlDTULICPTER ia RSN

OFTERf A DIVSION AL RIGHnS RESolD

T

A Suidauy t tton It, o LTD

2i

23 JUNE 1994

TEMPORARY REVISION - 8 AUGUST 1997 CANCELED - 5 SEPTEMBER 2000

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4

ROTOR

CRAFT

FLIGHT

SIN 36067

MANUAL

AND

SUBSEQUENT

TEMPORARY

MAST

TORQUE

SIGNAL

ASB

THlS MANUAL

SHALL

REVISION

CONDITIONER

412-97-90

SE IN THE HELICOPTER

DURING

ALL OPERATIONS

Bell

TEMPORARY REVISION - 8 AUGUST 1997 REVlSlON 1 - 18 DECEMBER 1997

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOG OF REVISIONS

Temporary Revision

1 ..........

8 AUG97

18 DEC 97

LOG OF PAGES

Page

Revision NO.

Page

Revision NO.

Temp/REV 1 Title A-B . . . . . . . . . . . . . . . . . . . . ..Temp/Rev 1 NOTE: Figures 4-4 through 4-8 are provided in their entirety for ease of handling 4-11 4-15 ..... Temporary 4-15A-4-15H Temporary 4-16 -4-18 (no change) Temporary 4-19 -4-30 Temporary 4-30A - 4-30M Temporary 4-31-4-36 ...................................TemplRevl

TEMPORARY

REVlSlON - 8 AUG 97 REVISION 1 -18 DEC 97

A

BHT-412-FM-4

LOG OF REVISIONS Original ....................... 0.................... 23 JUN 94 Revision ...................... 1.................... 23 DEC 94 Revision ...................... 2.................... 06 JUN 95 Revision ...................... 3.................... 23 JUN 95 Revision ...................... 4.................... 29 AUG 95 Revision ...................... 5.................... 26 OCT 95 Revision ...................... 6.................... 23 FEB 96 Revision ...................... 7.................... 22 AUG 96 Revision ...................... 8.................... 10 SEP 96 Revision ...................... 9................... 12 MAR 97 Revision ..................... 10.................. 12 MAR 97 Revision ..................... 11.................... 23 JUL 97 Revision ..................... 12................... 08 AUG 97

Revision......................13 .................. 14 NOV 97 Revision......................14 .................. 14 NOV 97 Revision......................15 .................. 26 MAR 98 Revision......................16 .................. 21 APR 98 Revision......................17 .................. 14 MAY 98 Revision......................18 .................. 28 OCT 98 Revision......................19 .................. 20 DEC 99 Revision......................20 .................. 26 MAY 00 Revision......................21 ................... 05 SEP 00 Revision......................22 ................... 11 JUN 01 Revision......................23 .................. 23 MAY 02 Revision......................24 .................. 05 NOV 02

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O LOG OF PAGES

PAGE

REVISION NO.

Cover........................................................24 Title...........................................................24 NP .............................................................24 A — B .......................................................24 C/D............................................................24 i ...................................................................0 ii ................................................................19 iii/iv .............................................................0 1-1.............................................................18 1-2.............................................................19 1-3.............................................................24 1-4.............................................................23 1-4A/1-4B .................................................18 1-5.............................................................20 1-6.............................................................22 1-7...............................................................9 1-8 — 1-10 ................................................17 1-11...........................................................23 1-12...........................................................19 1-13 — 1-16 ................................................5 1-17...........................................................20 1-18 — 1-21 ................................................5 1-22...........................................................17

REVISION NO.

PAGE

1-23 — 1-24 ............................................. 10 2-1 — 2-2 ................................................. 19 2-3 .............................................................. 0 2-4 ............................................................ 15 2-5 ............................................................ 20 2-6 ............................................................ 17 2-7 ............................................................ 19 2-8 .............................................................. 0 2-9 — 2-12 ............................................... 17 2-12A/2-12B ............................................. 20 2-13 .......................................................... 17 2-14 .......................................................... 22 2-15 ............................................................ 9 2-16 .......................................................... 19 2-17 .......................................................... 20 2-18 .......................................................... 19 2-19 .......................................................... 23 2-20 .......................................................... 20 2-20A/2-20B ............................................. 19 2-21 .......................................................... 15 2-22 .......................................................... 13 2-23/2-24 .................................................... 0 3-1 — 3-2 ................................................. 19

NOTE Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. 05 NOV 2002

Rev. 24

A

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FM-4

LOG OF APPROVED REVISIONS

8 AUG 97 Temporary ......................... Revision 1 .......................... 18 DEC 97 Canceled .......................... 05 SEP 00

NOTICE

This Temporary Revision has been cancelled. Subject data has been incorporated in BHT-412-FM-4 at Revision 21. Remove and discard all Temporary pages dated 8 August 1997 from this manual including these canceled pages.

APPROVED:

SEP

52000

MANAGER

ROTORRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170 B

TEMPORARY REVISION - 8 AUG 97 CANCELED-5 SEP 2000



Temporary ...................... Revision 1 ..........................

.8 AUG 97 18 DEC 97

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170 B

TEMPORARY REVISION 8 AUG 97 REVISION 1 - 18 DEC 97

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOG OF FAA APPROVED REVISIONS

Original ....................... 0.................... 23 JUN 94 Revision ...................... 1.................... 23 DEC 94 Revision ...................... 2.................... 06 JUN 95 Revision ...................... 3.................... 23 JUN 95 Revision ...................... 4....................29 AUG 95 Revision ...................... 5.................... 26 OCT 95 Revision ...................... 6.................... 23 FEB 96 Revision ...................... 7....................22 AUG 96 Revision ...................... 8.................... 10 SEP 96 Revision ...................... 9................... 12 MAR 97 Revision ..................... 10.................. 12 MAR 97 Revision ..................... 11....................23 JUL 97 Revision ..................... 12...................08 AUG 97

APPROVED:

Revision......................13.................. 14 NOV Revision......................14.................. 14 NOV Revision......................15 .................. 26 MAR Revision......................16.................. 21 APR Revision...................... 17.................. 14 MAY Revision......................18.................. 28 OCT Revision...................... 19.................. 20 DEC Revision......................20.................. 26 MAY Revision......................21 ................... 05 SEP Revision......................22 ................... 11 JUN Revision.....................23.................. 23 MAY Revision......................24.................. 05 NOV

DATE:

NOV

97 97 98 98 98 98 99 00 00 01 02 02

5 2002

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

05 NOV2002

Rev. 24

C/D

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FM-4

GENERAL INFORMATION

ORGANIZATION

The Rotorcraft Flight Manual is divided into five sections as follows: Section 1

-

LIMITATIONS

Section 2

-

NORMAL PROCEDURES

Section 3

-

EMERGENCY AND MALFUNCTION PROCEDURES

Section 4

-

PERFORMANCE

Section 5

-

WEIGHT AND BALANCE

Appendix A

-

OPTIONAL EQUIPMENT

Sections 1 through 4 contain the FAA approved data necessary to operate the basic helicopter in a safe and efficient manner. Section 5 contains Weight and Balance data. Appendix A contains the listing of approved Supplements for optional equipment which shall be used in conjunction with the basic Flight Manual when respective optional equipment kits are installed. The Manufacturer's Data manual, (BHT412-MD-4) contains additional information to be used in conjunction with the Flight Manual.

TERMINOLOGY

WARNINGS, CAUTIONS, AND NOTES ARE USED THROUGHOUT THIS MANUAL TO EMPHASIZE IMPORTANT AND CRITICAL INSTRUCTIONS AS FOLLOWS:

AN OPERATING PROCEDURE, PRACTICE, ETC., WHICH, IF NOT CORRECTLY FOLLOWED, COULD RESULT IN PERSONAL INJURY OR LOSS OF LIFE. . ................. CAUTION

AN OPERATING PROCEDURE, PRACTICE, ETC., WHICH, IF NOT STRICTLY OBSERVED, COULD RESULT IN DAMAGE TO OR DESTRUCTION OF EQUIPMENT. NOTE

AN OPERATING PROCEDURE, ETC., WHICH IS CONDITION, ESSENTIAL TO HIGHLIGHT.

WORDS USE OF PROCEDURAL

The concept of procedural word usage and intended meaning which has been adhered to in preparing this manual is as follows: "Shall" has been used only when application of a procedure is mandatory. "Should" has been used only application of a procedure recommended.

when is

"May" and "need not" have been used only is of a procedure when application optional. "Will" has been used only to indicate futurity, never to indicate a mandatory procedure.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4

I

ABBREVIATIONS

DC

- Direct Current

Abbreviations used through out this manual are defined as follows:

DCPL

-Decouple

DECEL

- Deceleration

DECR

- Decrease

DEG

- Degrees

DH

-Decision Height

DET

- Detector

EEPROM

- Electrically Erasable ProgramableRead Only Memory

EMERG

-Emergency

ENC

- Encoding

F

- Fahrenheit

FT FUELTRANS GEN GOV

-

AC

- Alternating Current

ADD

- Air DataDisplay

ADF

- Automatic Direction Finder

ADI

- Attitude Director Indicator

AFCS

- Automatic Flight Control System

AHRS

- Attitude Headingand ReferenceSystem

ALTN

- Alternate

AOG

- Aircraft on Ground

API

- Actuator Position Indicator

AP1/AP2

- Autopilot 1/ Autopilot 2

ATTor ATTD

- Attitude

AUX

-Auxiliary

BAT

- Battery

BLWR

- Blower

HD HDG Hp

C

- Celsius

HSI

CDI

- Course Deviation Indicator

CCP

- CompassControl Panel

HYDR IAS

- Hydraulic - Indicated Airspeed

CG

- Centerof Gravity

ICS

CPL

-Couple

CPLT

- Copilot

CRS

- Course

IFR IGE IMC

CYCCTR

- Cyclic Center

- IntercommunicationSystem - Instrument Flight Rules - In Ground Effect - Instrument Meteorological Conditions - Increase - Interconnect

II

Rev. 19

INCR INTCON

Force Trim or foot/feet Fuel Transfer Generator Governor Density Altitude

- Heading - Pressure Altitude

- Horizontal Situation Indicator


BHT-412-FM-4

INV

-

Inverter

PNL

-

Panel

ITT

-

Interturbine

PWR

-

Power

PSI

-

Instantaneous Vertical

-

Pounds per Square Inch

-

Knots Calibrated

RAD ALT

-

Radar Altimeter

Airspeed

REL

-

Release

-

Knots Indicated

RH

-

Right Hand

LE

-

Level

SAS

-

Stability Augmentation System

LF

-

Low Frequency

SBY ATT

-

Standby Attitude

LH

-

Left Hand

SL

-

Sea Level

LRC

-

Long Range Cruise

STBY

-

Standby

LT

-

Light

SYNC

-

Synchronize

MAG

-

Magnetic

SYS

-

System

MAX END

-

Maximum Endurance

VAC

-

MCP

-

Maximum Continuous

Volts Alternating Current

Power

VDC

-

Volts Direct Current

IVSI

KCAS KIAS

Temperature

Speed Indicator

Airspeed

N,

-

Gas Producer RPM

VEL HLD

-

Velocity Hold

Nil

-

Power Turbine RPM

VFR

-

Visual Flight Rules

NAV

-

Navigation

VG

-

Vertical Gyro

NON-ESNTL

-

Non-essential

VLF

-

Very Low Frequency

NORM

-

Normal

VMC

-

NR

-

Rotor RPM

Visual Meteorological Conditions

OAT

-

Outside Air

VNE

-

Never Exceed Speed

Temperature

VTOCS

-

One Engine Inoperative

Takeoff Climbout Speed

XPDR

-

Transponder

OEI

-

OGE

-

Out of Ground Effect

XFEED

-

Crossfeed

OVRD

-

Override

XMSN

-

Transmission

PART SEP

-

Particle Separator iii/iv

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FM-4

TABLE OF CONTENTS

Subject

INTRODUCTION ........................................................................ BASIS OF CERTIFICATION ........................................ ......... TYPES OF OPERATION............................................................. PASSENGER ............... ......................... CARGO ................................................................................... FLIGHT CREW ......................................................................... CONFIGURATION....................................................................... REQUIRED EQUIPMENT........................................................ OPTIONAL EQUIPMENT........................................................ DOORS OPEN/REMOVED ........................................ WEIGHT AND CENTER OF GRAVITY ....................................... WEIGHT................................................................................... CENTER OF GRAVITY .......................................................... AIRSPEED................................................................................... ALTITUDE................................................................................... MANEUVERING.......................................................................... PROHIBITED MANUEVERS................................................. CLIMB AND DESCENT........................................................... SLOPE LANDING ........................................ ......... HEIGHT - VELOCITY ........................................ AMBIENT TEMPERATURES...................................................... ELECTRICAL.............................................................................. BATTERY................................................................................ GENERATOR ............... ...................... ... STARTER .............................................................................. GROUND POWER UNIT ........................................................ POWER PLANT .......................................................................... GAS PRODUCER RPM........................................................... POWER TURBINE RPM (Nl) ................................................. INTERTURBINE TEMPERATURE.......................................... ENGINE TORQUE................................................................... FUEL PRESSURE................................................................... ENGINE OIL PRESSURE................................................. ENGINE OIL TEMPERATURE................................................ COMBINING GEARBOX OIL PRESSURE............................. COMBINING GEARBOX OIL TEMPERATURE .....................

TRANSMISSION ........................................

Page Number

Paragraph

.........

TRANSMISSION OIL PRESSURE.......................................... TRANSMISSION OIL TEMPERATURE .................................. MAST TORQUE ........................................ ......

1-1.................... 1-2.................... 1-3.................... .............. 1-3-A 1-3-B .............. 1-4.................... 1-5.................... 1-5-A .............. 1-5-B .............. 1-5-C.............. 1-6.................... 1-6-A.............. 1-6-B.............. 1-7.................... 1-8.................... 1-9.................... 1-9-A .............. 1-9-B .............. 1-9-C .............. 1-10.................. 1-11.................. 1-12.................. 1-12-A ......... 1-12-B............ 1-12-C ............ 1-12-D ............ 1-13.................. 1-13-A ............ 1-13-B ............ 1-13-C ............ 1-13-D............ 1-13-E............ 1-13-F ............ 1-13-G............ 1-13-H............ 1-13-J.............

1-3 1-3 1-3 1-3 1-3 1-4 1-4 1-4 1-4 1-4A/1-4B 1-4A/1-4B 1-4A/1-4B 1-4A/1-4B 1-5 1-5 1-6 1-6 1-6 1-6 1-6 1-6 1-6 1-6 1-6 1-7 1-7 1-7 1-7 1-7 1-8 1-8 1-8 1-8 1-9 1-9 1-9

1-14.................. 1-14-A............ 1-14-B............ 1-14-C............

1-9 1-9 1-9 1-10

Rev.18

-11


BHT-412-FM-4

TABLE OF CONTENTS (CONT)

Subject

Paragraph

ROTOR....................................................................................... ROTOR RPM- POWER ON ................................................ ROTOR RPM - POWER OFF .............................................. ROTOR RPM- GROUND OPERATION ............................... HYDRAULIC................................................................................ HYDRAULIC PRESSURE....................................................... HYDRAULIC TEMPERATURE ............................................... FUEL AND OIL......................................... FUEL ....................................................................................... OIL....................................................................................... ROTOR BRAKE ............................................. LANDING GEAR ......................................................................... INSTRUMENT MARKINGS AND PLACARDS........................... HEATER...................................................................................... HOIST PENALTY REGION......................................................... AHRS ALIGNMENT ....................................................................

Page Number 1-10 1-10 1-10 1-10 1-10 1-11 1-11 1-11 1-11 1-11 1-11 1-12 1-12 1-12 1-12 1-12

1-15 ........... 1-15-A............ 1-15-B............ 1-15-C............ 1-1 ................ 1-16-A............ 1-16-B............ 1-17 ................. 1-17-A............ 1-17-B............ 1-18................ 1-19................ 1-20................ 1-21................ 1-22................ 1-23................

LIST OF FIGURES

Figure Number

Title

Gross weight center of gravity charts ..................................... Weight-altitude-temperature limitations for takeoff, landing, and in-ground-effect maneuvers ............................... Maximum speed - sideward and rearward flight, crosswind and tallwind at a hover ........................................... Height-velocity diagram (OEI) ................................................ Instrument markings ................................................................. Placards and decals.................................................................. Hoist C.G. envelope...................................................................

1-2

Rev. 19

Page Number

1-1..................

1-13

1-2..................

1-15

1-3.................. 1-4.................. 1-5.................. 1-6.................. 1-7......... 1-.........

1-16 1-17 1-18 1-23 1-24

FAA APPROVED

BHT-412-FM-4

!ection 1 *IMITATION!

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O 1-1.

INTRODUCTION

Compliance with limitations in this section is required by appropriate operating rules.

responsibility to ensure helicopter is properly loaded so entire flight is conducted within limits of gross weight center of gravity charts (Figure 1-1).

1-2.

1-3-B.

BASIS OF CERTIFICATION

This helicopter is certified under FAR Part 29, Category A and B. Refer to BHT-412-FMS-62.4 for Category A Operations.

1-3.

TYPES OF OPERATION

Basic configured helicopter is approved as a fifteen-place helicopter and is certified for operation under day or night VFR and IFR nonicing conditions.

1-3-A.

PASSENGER

NOTE Refer to Section 5 for loading tables to be used in weight/CG computations.

Outboard facing seats should not be occupied unless at least four forward or aft facing passenger seats are occupied.

Above loading does not apply if cargo or a combination of cargo and passengers are being transpor ted. It shall then be pilot

CARGO

Maximum allowable deck loading for cargo is 100 pounds per square foot (4.9 kg/100 sq cm). Deck mounted cargo tiedown fittings are provided and have an airframe structural capacity of 1250 pounds (567.0 kilograms) vertical and 500 pounds (226.8 kilograms) horizontal per fitting. Provisions for installation of cargo tiedown fittings are incorporated in aft cabin bulkhead and transmission support structure and have an airframe structural capacity of 1250 pounds (567.0 kilograms) at 90 degrees to bulkhead and 500 pounds (226.8 kilograms) in any direction parallel to bulkhead. Cargo shall be secured by an approved restraint method that will not impede access to cargo in event of emergency.

NOTE Refer to the applicable operating rules for internal cargo operations.

Baggage compartment has a load limit of 400 pounds (181 kilograms), not to exceed 100 pounds per square foot (4.9 kg/100 sq cm). Refer to Section 5.

05 NOV 2002

Rev. 24

1-3

BHT-412-FM-4

1-4.

FAA APPROVED

FLIGHT CREW

The minimum flight crew consists of one pilot who shall operate helicopter from right crew seat. Refer to Section 5. for minimum crew station weight.

Two navigation receivers with auxiliary equipment appropriate to intended IFR route of flight. DME equipment.

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O

T h e l e f t c r e w s e a t m ay b e u s e d fo r a n additional pilot when approved dual controls are installed.

1-5.

CONFIGURATION

1-5-A.

1-5-A-1.

REQUIRED EQUIPMENT AFCS

AFCS shall be disengaged or operated in SAS mode during prolonged ground operation, except as required for AFCS check.

ATC transponder.

Marker beacon receiver.

3 inch standby attitude indicator. Pilot IVSI.

Force trim.

Roof window blackout curtains.

EMERGENCY COMM panel, (if installed) (single pilot only).

AHRS shall be certified only on helicopters S/N 36248 and subsequent.

Digital AFCS preflight test (level 1 minimum) shall be accomplished prior to first flight of the day or before planned flight into IMC.

1-5-A-2.

1-5-B.

REQUIRED EQUIPMENT — IFR

Both autopilots, AP 1 and AP 2, shall be engaged in ATT mode during flight. Heated pitot-static system. Pilot windshield wiper.

Two VHF communications radios.

1-4

Rev. 23

23 MAY 2002

OPTIONAL EQUIPMENT

Refer to appropriate Flight Manual Supplement(s) for additional limitations, procedures, and performance data with optional equipment installed. See Appendix A. When a CVR is required by the operating rules, only headsets with a dynamic microphone are approved. CVR operation with other types of microphones requires FAA approval.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAAAPPROVED

1-5-C.

DOORSOPEN/REMOVED

Helicopter may be flown with doors open or removedonly with Bell StandardInterior(412705-501)or Bell DeluxeInterior (412-705-500) installed. Flight operation is approved for following alternative configurations during VFRconditions only: Symmetricalconfigurations:

Both crew doors removed.

Both sliding doors lockedopen or removedwith both hingedpanels installedor removed.

Asymmetricalconfigurations:

Cargo dooors can be openedor closed asymmetrically,to a lockedposition, with following restrictions: 1. Twoway communicationsbetween pilot and cabin crew member.

2. All crew membersand passengersare securedwith an approvedrestraint.

NOTE Opening or removing doors shifts helicopter center of gravity and reducesVNE.Referto Section5 and to Airspeedlimitations.

BHT-412-FM-4

1-6.

1-6-A.

WEIGHT AND CENTER OF GRAVITY WEIGHT

Maximum gross weight for takeoff and landing is 11,900pounds (5398kilograms).

Refer to Weight-Altitude-Temperature Limitations chart (figure 1-2) for maximum allowableweight for takeoff, landing, and IGE hover operation. Minimum gross weight for flight is 6400 pounds (2903kilograms).

Minimum combined crew weight at fuselage station 47.0 is 170pounds (77.1kilograms). 1-6-B.

1-6-B-1.

CENTEROF GRAVITY

LONGITUDINAL CENTER OF GRAVITY

Longitudinal centerof gravity limits vary from station 130 to 144, depending on gross weight. Refer to Gross weight center of gravity charts (figure 1-1). 1-6-B-2.

LATERALCENTEROF GRAVITY

Lateral center of gravity limits are 4.5 Inches (114.3millimeters) left and right of fuselage centerlinefor all gross weights.

Rev.18

1-4A/1-4B

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4

FAA APPROVED

1-7.

AIRSPEED

Minimum IFR airspeed is 60 KIAS.

Basic VNE is 140 KIAS from sea level to 3000 feet density altitude at all gross weights. VNE decreases for ambient conditions in accordance with airspeed limitations placard (paragraph 1-20). 1(paragraph 20).Maximum Airspeed shall not exceed 105 KIAS (or placarded VNE, if less) when operating above 81% mast torque. VNE with only one autopilot engaged is 115 KIAS (or placarded VNE, if less). If both autopilots are disengaged, basic VNE applies. VNE for steady state autorotation is:

105 KIAS at or below 10,000 feet (3048 meters) pressure altitude;

80 KIAS above 10,000 feet (3048 meters) pressure altitude.

VNE with cargo doors opened asymmetrically is 80 KIAS.

VNE with cargo doors in transit or in an unlocked position Is 60 KIAS.

VNE with doors symmetrically open or removed is 100 KIAS with Bell Helicopter installed Blanket Interior (412-705-501) or Deluxe Interior (412-705-500).

Maximum allowable airspeed for sideward or rearward flight at or below 3000 feet HD is 35 knots. See figure 1-3 for additional limitations. allowable tailwind or crosswind speeds for hover operations at or below 3000 feet HD is 35 knots. See figure 1-3 for additional limitations. Refer to Critical relative wind azimuths diagram in Section 4.

1-8.

ALTITUDE

NOTE NOTE

Refer to applicable operating rules for high altitude oxygen requirements.

Maximum operating pressure altitude is 20,000 feet (6096 meters).

Maximum density altitude for takeoff, landing, and in-ground-effect maneuvers is 14,000 feet (4267 meters). Refer to Weight-AltitudeTemperature Limitations chart (figure 1-2). Above 15,000 feet (4,572 meters) pressure altitude, restart shall be attempted in manual fuel control mode only.

VNE with doors symmetrically open or removed is 60 KIAS with Bell Helicopter installed energy attenuating passenger seats (412-706-002). NOTE Asymmetric door configuration is not authorized with energy attenuating seats installed.

Rev. 20

1-5

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FM-4

Below 15,000 feet (4,572 meters) pressure altitude, restart may be attempted In either manual or automatic fuel control mode.

1-9.

Minimum ambient temperature for operation at all altitudes is -40°C (-40°F).

MANEUVERING

PROHIBITED MANEUVERS

1-9-A.

NOTE

Refer to paragraph 1-13-F, Engine

Oil Pressure.

Aerobatic maneuvers are prohibited.

1-9-B. CLIMB AND DESCENT

1-12.

Maximum IFR rate of climb or descent is 1000

1-12-A.

feet per minute.

Maximum battery case temperature is 54.5°C (130°F), as Indicated by illumination of BATTERY TEMP warning light.

Maximum IFR approach slope is 5 degrees.

1-9-C.

SLOPE LANDING

Slope landings are limited to side slopes not to exceed 10 degrees.

1-10.

HEIGHT-VELOCITY

The height-velocity limitations are critical in the event of single engine failure during takeoff, landing, or other operation near the surface (figure 1-4). The AVOID area of the Height-Velocity diagram defines the combinations of airspeed and height above ground from which a safe single engine landing on a smooth, level, firm surface cannot be assured.

ELECTRICAL

BATTERY

WARNING

WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER

ILLUMINATION

OF

BATTERY

TEMP LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURN TO SERVICE.

Minimum ambient temperature for battery start when battery and helicopter have been cold soaked is -25°C (-13F).

1-12-B.

GENERATOR

The H-V diagram Is valid only when the Weight-Altitude-Temperature limitations are Continuous operation 0 to 75 amps not exceeded (figure 1-2). The diagram does Caution 75 to not define the conditions which assure Caution 75 to 150 amps continued flight following an engine failure NOTE nor the conditions from which a safe power off landing can be made. During OEI operation electrical loads may have to be reduced to remain below maximum 1-11. AMBIENT TEMPERA-

TURES

The maximum sea level ambient air temperature for operation is +51.7°C (+125°F) and decreases with pressure altitude at the standard lapse rate of 2°C (3.6°F)/1000 feet (305 meters) to 20.000 feet (6096 meters). 1-6

Rev. 22 11 JUN 2001

continuous limits.

BHT-412-FM-4

FAA APPROVED

T N I R I P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Maximum continuous

150 amps (each)

helicopter logbook detailing the extent of operation In excess of twin engine takeoff power limits. This does not apply to approved ITT limits for starting.

NOTE

Ammeter needle may deflect full scale momentarily during generator assisted start of second engine.

1-12-C. STARTER

Starter energizing times shall be limited as

1-13-A. GAS PRODUCER RPM (N1) 1-13-A-1. TWIN ENGINE OPERATION Continuous operation

61 to 103.1%

Maximum continuous

103.1%

Maximum for takeoff

103.1%

follows:

30 seconds ON

60 seconds OFF 30 seconds ON 5 minutes OFF

1-13-A-2. ONE ENGINE INOPERATIVE (OEI)

30 seconds ON

Continuous OEI

103.7

15 minutes OFF

2

103.7 to 109.2%

1-12-D. GROUND POWER UNIT

28 vdc ground power units for starting shall be limited to 1000 amps maximum.

1-13.

1/2

minute OEI

Maximum OEI

109.2%

1-13-B. POWER TURBINE RPM (Ni.)

POWER PLANT

Pratt and Whitney Aircraft of Canada, Ltd. PT6T-3D. NOTE

in 2'/2 minute or Operation Continuous OEI range is intended for emergency use only, when one engine becomes inoperative due to an actual malfunction.

Anytime an engine is operated in an OEI range, an entry shall be made in the

Minimum in cruise

97%

Minimum for hover, takeoff, and climb

100%

Maximum continuous

100%

Operation with ENG torque at or below 30% 100 to 104.5% Maximum with ENG torque at or below 30% 104.5% Rev.9

1-7

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 1-13-C. INTERTURBINE TEMPERATURE

NOTE Either ITT gage shall be installed in pairs.

1-13-C-1. TWIN ENGINE OPERATION

GAGE PN 212-075-067-115 Continuous 300 to 810 °C Maximum continuous 810 °C Maximum transient (5 seconds maximum above 810° ) 925 °C Maximum for starting (2 seconds maximum ° above 960 )

1090 °C

GAGE PN 212-075-067-119 Continuous 300 to 810 °C Maximum continuous 810 °C Maximum transient seconds maximum above 810 ° )

(5

940 °C

Maximum for starting (2 seconds maximum above 960 ° ) 1090 °C

CAUTION

INTENTIONAL USE OF ITT ABOVE 810°C IS PROHIBITED DURING NORMAL OPERATIONS EXCEPT DURING START.

1-13-C-2. ONE ENGINE INOPERATIVE (OEI) GAGE PN 212-075-067-115

2 1/2 minute OEI Maximum OEI

820 to 940 °C 940 °C

1-13-D. ENGINE TORQUE

ONE ENGINE INOPERATIVE (OEI) Continuous OEI

5 to 73.2%

Maximum Continuous OEI 21/2 minute OEI Maximum OEI

73.2% 73.2 to 81% 81%

1-13-E. FUEL PRESSURE

Minimum Continuous Maximum

operation

4 psi 4 to 35 psi 35 psi

1-13-F. ENGINE OIL PRESSURE NOTE During extremely cold ambient temperatures, Idle rpm will be high and the ENGINE OIL pressure may exceed maximum limits for up to two minutes after starting. NOTE Either Engine Oil Pressure/ Temperature gage shall be installed in pairs. Refer to paragraph 1-11, AMBIENT AIR TEMPERATURE.

GAGE PN 209-070-262-113 Minimum for flight Idle Operation below 79% N, rpm Continuous operation Maximum

40 psi

40 to 90 psi 90 to 115 psi 115 psi

Maximum continuous OEI

820 °C

GAGE PN 209-070-262-109

2 1/2 minute OEI Maximum OEI

820 to 925 °C 925 °C

Minimum for flight idle Operation below 79% N, rpm Continuous operation Maximum

GAGE PN 212-075-067-119 Maximum continuous OEI 1-8

Rev. 17

820 °C

40 psi

40 to 80 psi 80 to 115 psi 115 psi

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Continuous operation

1-13-G. ENGINEOIL TEMPERATURE

0 to 115 °C

Maximum for MIL-L-7808 oil 115 ° C

Minimum

0 °C

Continuous operation

0 to 115 °C

Maximum for MIL-L-23699 oil 120 °C Maximum for DOD-L85734oil

120 °C

Maximum for MIL-L-7808 oil 115 °C Maximum for MIL-L-23699 oil 120 °C Maximum for DOD-L85734 oil

1-14. TRANSMISSION

1-14-A. TRANSMISSION OIL PRESSURE

120 °C

NOTE

1-13-H. COMBININGGEARBOXOIL PRESSURE

During extremely cold ambient temperatures, idle rpm will be high and the XMSN OIL pressure may exceed maximum limits for up to two minutes after starting..

NOTE

During extremely cold ambient temperatures,GEAR BOX OIL pressure may exceed maximum limits for up to two minutes after starting.

Minimum for idle

Minimum for idle

30 psi

Idle range

30 to 40 psi


40 to 70 psi

Maximum

70 psi

40 psi

Operation below 94% N11rpm

40 to 60 psi


60 to 80 psi

Maximum

80 psi

1-14-B. TRANSMISSION OIL TEMPERATURE Continuous operation

15 to 110 °C

Maximum

110 °C

1-13-J. COMBININGGEARBOXOIL TEMPERATURE Minimum

0 °C Rev.17

1-9

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 1-14-C. MAST TORQUE

1-14-C-1. TWIN ENGINE OPERATION Maximum continuous

100%

Operation with ENG torque at or below 30% 100 to 104.5% Maximum with ENG torque at or below 30% 104.5%

81%

Takeoff range (5 minutes maximum) 81 to 100% Maximum

Maximum continuous

1-15-B. ROTOR RPM - POWER OFF

100%

Minimum for autorotation with gross weight below 8000 pounds (3629 kg) 80% Power off operation with

gross weight below 8000

TAKEOFF POWER SHALL NOT BE USED ABOVE 105 KIAS.

pounds (3629 kg)

Minimum for autorotation with gross weight at or

above 8000 pounds (3629

CAUTION

80 to 104.5%

kg)

91%

Maximum

104.5%

WHEN OPERATING NEAR THE MAXIMUM MAST TORQUE LIMIT,

INADVERTENT OVERTORQUE MAY OCCUR DURING MANEUVERING FLIGHT

1-15-C. ROTOR RPM - GROUND OPERATION

CONDITIONS INVOLVING TURNS

AND/OR NOSE DOWN ATTITUDE CHANGES. DECREASE POWER TO 90% MAST TORQUE PRIOR TO

MANEUVERINGHELICOPTER.

Intentional use of mast torque over 100% is prohibited.

Minimum

77%

Minimum with stick

centering indicator system inoperative

97%

Transient (avoid steady state operations) 26 to 77%

1-16. HYDRAULIC

1-15. ROTOR

1-15-A. ROTOR RPM - POWER

ON

NOTE

Refer

to

BHT-412-MD-4

approved fluids and vendors.

Minimum

97%


97 to 100%

for

Hydraulic fluid type MIL-H-5606 (NATO H515) shall be used at all ambient

temperatures.

Both hydraulic systems shall be operative

prior to takeoff. 1-10

Rev.17

FAA APPROVED

BHT-412-FM-4

WARNING

1-17-B. 1-17-B-1.

THE HELICOPTER IS NOT CONTROLLABLE WITH BOTH HYDRAULIC BOOST SYSTEMS INOPERATIVE.

OIL ENGINE AND COMBINING GEARBOX

Oil conforming to PWA Specification No. 521 Type I and MIL-L-7808 (NATO 0-148) may be used at all ambient temperatures. Oil conforming to PWA Specification No. 521 Type Il and MIL-L-23699 (NATO 0-156), or DOD-L-85734AS limited to ambient temperatures above -40°C (-40°F).

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O 1-16-A.

HYDRAULIC PRESSURE

Minimum

600 psi

Caution

600 to 900 psi


900 to 1100 psi

Maximum

1100 psi

1-16-B.

HYDRAULIC TEMPERATURE

Maximum

1-17-B-2.

Oil conforming to DOD-L-85734AS (Turbine Oil 555) is the only approved oil for use in the transmission and gearboxes.

88°C

1-18.

1-17.

FUEL AND OIL

NOTE Refer to BHT-412-MD-4 for lists of approved fuels, oils, capacities, and vendors.

1-17-A.

TRANSMISSION, INTERMEDIATE AND TAIL ROTOR GEARBOX

ROTOR BRAKE

Engine starts with rotor brake engaged are prohibited. Rotor brake application is limited to ground operation and shall not be applied until both engines are shut down and rotor rpm has decreased to 40% NR or below.

FUEL

Fuel conforming to ASTM D-1655 Type B, NATO F-40, or MIL-T-5624 Grade JP-4 may be used at all ambient temperatures.

Fuel conforming to ASTM D-1655 Type A or A-1, NATO F-44, MIL-T-5624 Grade JP-5, NATO F-34, or MIL-T-83133 Grade JP-8, limited to ambient temperatures above -30°C (-22°F).

23 MAY 2002

Rev. 23

1-11


BHT-412-FM-4

1-19.

LANDING GEAR

for Bell Helicopter approved Hoists.

No Flight Manual Limitations.

1-20.

AND PLACARDS

NOTE Illustrations shown in figure 1-5 are artist representations and may or may not depict actual approved instruments due to printing limitations. Instrument operating ranges and limits shall agree with those presented in this section.

Refer to Figure 1-5 for instrument range markings and Figure 1-6 for placards and decals.

1-21.

THIS PENALTY REGION IS VALID FOR ALL HOIST INSTALLATIONS.

OPERATION IN PENALTY REGION AIRWORTHINESS AFFECTS ROTOR OF LIMITATIONS COMPONENTS (REFER TO BHT-412MM).

1-23.

ALIGNMENT AHRS ALIGNMENT

AHRS

To perform in-flight/shipboard

alignment, the following conditions must be met:

HEATER

Heater shall not be operated when OAT is above 21°C (69.8°F).

1-22.

WARNING

INSTRUMENT MARKINGS

HOIST PENALTY REGION

Pilot shall know C.G. at time of hoist operation to determine if C.G. is within penalty region of figure 1-7, Hoist C.G. envelope.

°

Pitch

Less than ± 10

Roll Yaw Rate

Less than ± 5 Less than 1°/sec

Lateral and Logitudinal

Less than 0.05g

Each hoist operation performed is defined as

Acceleration

an extension and retraction of hoist cable while hovering with any weight attached.

For a minimum of 30 seconds for attitude and

Refer to BHT-412-FMS-7 or BHT-412-FMS-26

1-12

Rev. 19

heading to become valid.


BHT-412-FM-4

ENGLISH UNITS

BHT-412EP-FM-1-1-1

Figure 1-1.

Gross weight center of gravity charts (Sheet 1 of 2).

Rev. 5

1-13

BHT-412-FM-4


359

3432

400

5398-

5200

/

5000-

--

4800

4400

-4200

3302

4000

3992

FOWARDLIMIT

AFT LIMIT

3800

3600 -

3400 I 3200 3000

MINIMUMWEIGHT

3250

3300

3350

3400

3450

3500

3658

3550

3600

650

LONGITUDINAL C.G. FUSELAGESTA.-mm

METRICUNITS

BHT-412EP-FM-1-1-2

Figure 1-1. Gross weight center of gravity charts (Sheet 2 of 2).

1-14

Rev.5


BHT-412-FM-4

NOTE: ALLOWABLEGROSSWEIGHTSOBTAINEDFROMTHIS CHART MAY EXCEED CONTINUOUSHOVERCAPABILITYUNDERCERTAINAMBIENT CONDITIONS. REFERTO HOVERCEILINGCHARTSIN SECTION4.

-40-20

0 20 OAT- C

40 60

9

4.0

10

11

4.5 5.0 GROSSWEIGHT

12 LBX1000

5.4

kg x 1000

BHT-412EP-FM-1-2

limitations for takeoff, landing, and in-groundFigure 1-2. Weight-altitude-temperature effect maneuvers. Rev. 5

1-15

BHT-412-FM-4


NOTE:

OGECRITICALRELATIVEWIND AZIMUTH

IGECRITICALRELATIVE WIND AZIMUTH

14,000 FT. DENSITY ALTITUDE

Figure 1-3.

Maximum speed -

Rev. 5

sideward and rearward flight, crosswind and tailwind at a hover.

BHT-412-FM-4


HEIGHT - VELOCITY DIAGRAM

FOR SMOOTH. LEVEL. FIRM SURFACES

120 NOTE: HELICOPTER CONFIGURATION AND ABOVE

400 AND ABOVE

SHALL COMPLYWITH THE WEIGHT .

LIMITS ALTITUDETEMPERATURE

375 380

AS PRESENTEDIN FIGURE 1-2FOR HEIGHT - VELOCITY DIAGRAM TO BE

360

VALID.

-114.3 -110

340

-100

320 300

280

80

260 220 200

60

180 160

0

140

40

120

10

30

80

-20

40

10

0

10

20

30

40

VNE

INDICATED AIRSPEED - KNOTS

BHT-412EP-FM-1-4

Figure 1-4. Height-velocity diagram (OEI).

Rev. 20

1-17

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O I

1-18 Rev. 5

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


412EP-F-1-5-2 Figure 1-5. Instrument markings (Sheet 2 of 5)

I Rev. 6

1-19

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4

FAA APPROVED

412EP-F-1-5-4 Figure 1-5. Instrument markings (Sheet 4 of 5)

I



BHT-412-FM-4

Rev. 10

1-231

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O :,.:':. ........

o

O

o

o

oC 0

—

rr, CO

T N I Section 2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4

FAA APPROVED

TABLE OF CONTENTS

Subject

Paragraph

Page Number

INTRODUCTION ......................................................................... FLIGHT PLANNING .................................................................... TAKEOFF AND LANDING ...................................................... WEIGHT AND BALANCE ....................................................... PREFLIGHT CHECK................................................................... BEFORE EXTERIOR CHECK................................................. EXTERIOR CHECK................................................................. INTERIOR AND PRESTART CHECK......................................... ENGINE START .......................................................................... ENGINE 1 START ................................................................... ENGINE 2 START ................................................................... POST START .......................................................................... ENGINE FAILS TO START..................................................... DRY M OTORING RUN............................................................ COLD W EATHER START....................................................... SYSTEMS CHECK...................................................................... STICK CENTERING INDICATOR CHECK............................. FORCE TRIM CHECK............................................................. PRELIMINARY HYDRAULIC CHECK .................................... ENGINE FUEL CONTROL CHECK........................................ GOVERNOR CHECK .............................................................. FUEL CROSSFEEDAND INTERCONNECT VALVE CHECK ELECTRICAL SYSTEMS CHECK.......................................... HYDRAULIC SYSTEMS CHECK............................................ AFCS CHECK ......................................................................... AHRS CHECK (S/N 36248 AND SUB) ................................... CABIN HEATER CHECK........................................................ BEFORE TAKEOFF.................................................................... POWER ASSURANCE CHECK.............................................. PROLONGED GROUND OPERATION .................................. OPERATION IN OR NEAR MAGNETIC DISTURBANCES (AHRS EQUIPPED SHIPS - S/N 36248 AND SUB)............. TAKEOFF.................................................................................... IN-FLIGHT OPERATIONS .......................................................... MANEUVERING WITH AFCS IN SAS MODE ........................ MANEUVERING WITH AFCS IN ATT MODE......................... DESCENT AND LANDING.......................................................... ENGINE SHUTDOWN.................................................................

2-1.................. 2-2.................. 2-2-A .............. 2-2-B .............. 2-3.................. 2-3-A .............. 2-3-B .............. 2-4.................. 2-5.................. 2-5-A .............. 2-5-B .............. 2-5-C .............. 2-5-D .............. 2-5-E .............. 2-5-F .............. 2-6 .................. 2-6-A .............. 2-6-B .............. 2-6-C .............. 2-6-D .............. 2-6-E .............. 2-6-F .............. 2-6-G .............. 2-6-H .............. 2-6-J............... 2-6-K .............. 2-6-L .............. 2-7.................. 2-7-A .............. 2-7-B ..............

2-3 2-3 2-3 2-3 2-4 2-4 2-4 2-7 2-9 2-9 2-11 2-11 2-12 2-12 2-12A/2-12B 2-13 2-13 2-13 2-13 2-14 2-14 2-14 2-15 2-15 2-16 2-17 2-18 2-18 2-19 2-19

2-7-C.............. 2-8 .................. 2-9.................. 2-9-A .............. 2-9-B .............. 2-10................ 2-11................

2-19 2-19 2-20 2-20 2-20 2-20 2-21

Rev. 19

2-1

FAA APPROVED


POSTFLIGHT CHECK ................................................................

2-12................

2-22

LIST OF FIGURES

Figure Number

Title

Preflight check sequence .....................................................

1

2-2

Rev. 19

2-1................

Page Number

2-23/2-24


BHT-412-FM-4

2-1. INTRODUCTION

2-2. FLIGHT PLANNING

This section contains instructions and procedures for operating the helicopter from the planning stage, through actual flight conditions, to securing the helicopter after landing.

Each flight should be planned adequately to ensure safe operations and to provide the pilot with the data to be used during flight. Essential weight, balance, and information should be performance compiled as follows:

Normal and standard conditions are assumed in these procedures. Pertinent data in other sections is referenced when applicable.

Check type of flight to be performed and destination.

The instructions and procedures contained herein are written for the purpose of standardization and are not applicable to all situations. The minimum, normal, maximum, cautionary operation ranges for helicopter and its subsystems indicated by instrument markings placards.

and the are and

Select appropriate performance charts to be used.

2-2-A. TAKEOFFAND LANDING

Refer to the Limitations section for takeoff and landing weight limits and to the Performance section for takeoff and landing distance information.

2-2-B. WEIGHTAND BALANCE

Anytime an operating limitation is exceeded, an appropriate entry shall be made in the helicopter logbook. The entry shall state which limit was exceeded, the duration of time, the extreme value attained, and any additional information essential in determining the maintenance action required.

Determine proper weight and balance of the helicopter as follows:

These instrument markings and placards represent careful aerodynamic calculations that are substantiated by flight test data.

Check that loading limitations listed in Section 1 have not been exceeded.

Refer to Section 1, Limitations, subsystems restrictions.

Consult Section 5 for instructions.

Compute takeoff and anticipated landing gross weight, check helicopter center of gravity (CG) locations, and determine weight of fuel, oil, payload, etc.

for

2-3

BHT-412-FM-4


2-3. PREFLIGHTCHECK

The pilot is responsible for determining whether the helicopter is in condition for safe flight. Refer to figure 2-1 for preflight check sequence. NOTE

The pilot walk-around and interior checks are outlined in the following procedures. The preflight check is not intended to be a detailed mechanical check, 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.

2-3-A. BEFOREEXTERIOR CHECK Flight planning - Completed. Gross weight and CG to Section 5). Publications -

Compute (refer

Portable fire extinguishers and security.

Condition

Drain samples as

FUEL TRANS switches -

BOOST PUMPswitches -

OFF.

OFF.

ENGINE 1 and ENGINE 2 FUEL switches - OFF. BAT BUS 1 switch 2-4

Rev.15

ON.

NOTE

If aft sumps fail to drain, sump valves may be operated manually.

Forward and middle fuel sumps samples as follows: Press-to-drain valves -

Drain

Press.

Fuel filters - Drain before first flight of day as follows: BOOST PUMPswitches -

ON.

ENGINE 1 and ENGINE 2 FUEL switches - ON. Fuel filter (left and right) samples.

-

Drain

ENGINE 1 and ENGINE 2 FUEL switches - OFF. BOOST PUMPswitches - OFF. BAT BUS 1 switch -

OFF.

Main and tail rotor blade tie downs Remove and stow. Pitot tube cover(s) stow.

Remove and

No. 1 and 2 engine air intake covers Remove and stow.

Checked.

Aft fuel sumps follows:

Aft fuel sump drain buttons (left and right) - Depress.

2-3-B. EXTERIORCHECK. WARNING

IF HELICOPTER HAS BEEN EXPOSED TO SNOW OR ICING CONDITIONS, SNOW AND ICE SHALL BE REMOVED PRIOR TO FLIGHT.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED 1.

2.

FUSELAGE -

BHT-412-FM-4

FRONT

N11governor spring - Check condition.

Cabin nose - Condition; all glass clean; wipers stowed.

Engine fire extinguisher - Check bottle pressure gauge and temperature range.

Remote hydraulic filter bypass indicator - Verify normal.

Combining gearbox filter - Check bypass indicator retracted.

Pitot tube(s) -

Oil cooler blower - Unobstructed.

Unobstructed.

Static ports (left and right) Unobstructed.

Avionics compartment - Security of components; check OVER TRQ warning

Cabin nose ventilators -

flag (cat's eye) for indication of overtorque. AFCS computers - Secured;

Unobstructed.

Nose compartment door - Secure.

compartment door secured.

Battery vent and drain tubes Unobstructed.

Access doors and engine cowling Secured.

Searchlight and landing light - Stowed.

Drain lines - Clean and unobstructed.

Antennas - Condition and security.

Engine exhaust ejectors - Covers removed; unobstructed.

FUSELAGE - CABINE LEFT SIDE.

Copilot door - Condition and operation; glass clean. Check security of emergency release handles.

Tailboom - Condition; access covers

Position lights - Condition.

secured.

Passenger door - Condition and operation; glass clean. Condition of pop-

Tail rotor driveshaft covers -

out windows.

CAUTION

Landing gear - Condition; handling wheels removed.

Passenger step (if Installed) - Condition and security.

3.

FUSELAGE -

Secured.

AFT LEFT SIDE

DO NOT BEND ELEVATOR TRAILING EDGE TAB.

Elevator - Condition and security. Check for spring condition by moving elevator toward the leading edge down

position.

No. 1 engine compartment -

Check.

No. 1 engine oil level - Verify actual presence of oil In sight gauge. Visually check oil level. Filler cap secured.

Tail rotor (90x) gearbox - Verify actual presence of oil in sight gauge. Visually check oil level, check filler cap and chip detector plug for security.

Rev. 20

2-5


BHT-412-FM-4

Tail rotor blades - Visually check condition and cleanliness.

Condition and

Passenger door-

Tail rotor - Condition and free movement on flapping axis.

operation, glass clean. Condition of pop-out windows.

Tail rotor yoke - Evidenceof static

Transmission oil -

stop contact damage (deformed static stop yield indicator). Tail skid -

Condition and security.

Intermediate (42°) gearbox -

Verify

Verify actual gauge.

presence of oil in sight Visually check oil level. Position lights - Condition.

Condition, handling

Landing gear -

actual presence of oil in sight gauge.

wheels removed.

cap and chip detector plug for

Passenger step (if installed)

Visually check oil level. Check filler security.

and security. Condition

Tailboom -

Condition.

5. FUSELAGE -

AFT RIGHT SIDE. Check compartment

Pilot door -

Condition and operation,

Check security of

glass clean.

Check.

-

Engine fire extinguisher

7. CABIN TOP Hub and sleeve assembly condition. Swashplate,

Check

bottle pressure gauge and

assembly,

support

Mast retaining nut -

Combining gearbox oil level - Verify actual presence of oil In sight gauge.

Yoke assembly -

Oil cooler blower -

Unobstructed.

No. 2 engine compartment No. 2 engine oil level -

Check.

Verify actual

and

Security and

Check general

Main rotor hub condition:

temperature range.

Visually checkoil level.

Check

collectivelever - Check condition. Main rotor pitch links condition.

unobstructed.

Tail rotor actuator -


emergency release handles.

Baggage compartment - Cargo secured, smoke detector condition, door secured.

Secured.

Condition.

Pitch horns -

Security

and

condition.

Elastomeric bearings, lead-lag dampers condition.

-

Check

general

presence of oil in sight gauge.

Blade retention bolts - Security

Access doors and engine cowling -

Droop restrainers - Security and condition.Verify droop stop clevis is

Visually check oil level. secured.

Filler cap

Secured.

Fuel filler secure cap. 2-6

CABIN RIGHTSIDE

6. FUSELAGE-

Rev. 17

Visually check quantity;

and proper latching.

in lower position of cam plate.

Simple pendulum absorbers Security and condition.

-


BHT-412-FM-4

Rotor blades - Visually check condition and cleanliness.

Main driveshaft and coupling Condition, security and grease leakage. Check Temp-Plates (four places each coupling) for evidence of elevated temperature indicated by dot changing color to black.

Protective breathing equipment (if installed) - Condition and properly serviced. NOTE

Opening or removing doors shifts helicopter center of gravity and reduces VNE. Refer to Section 1 and Section 5. Passenger doors - Secured.

CAUTION

Seat and pedals - Adjust.

Seat belt and shoulder harness - Fasten and adjust.

IF ANY TEMP-PLATE IS MISSING OR HAS BLACK DOTS, MAINTENANCE PERSONNEL SHALL ASSIST IN DETERMININGAIRWORTHINESS.

Shoulder harness inertia reel and lock Check.

Transmission oil filler cap - Secured.

No. 1 and No. 2 hydraulic reservoirs Visually check fluid levels; caps secured. Antenna(s) -


Combining Secured.

gearbox

Anticollision security.

light

oil filler

-

cap -

Condition

and

No. 1 and No. 2 engine air intakes Unobstructed; check particle separator doors closed. Engine and transmission Secured. Fresh air inlet screen -

cowling

-

Unobstructed.

Rotor brake reservoir cap - Security.

2-4.

INTERIOR AND PRESTART CHECK

Cabin interior - Cleanliness and security of equipment. Cargo and baggage (if applicable) Check security.

Directional control pedals Check freedom of movement, position for engine start. Flight controls friction as desired.

Position

EEPROM READ switch down position.

for start,

Locked

in

Transmission chip detector indicators Check, reset if required. Lower pedestal circuit breakers -

IN.

Collective control head switches - OFF. COMPASS CONTROL panel switches MAG position (normal slaved).

AHRS control panel switches - MAG position (locked). (S/N 36248 and subsequent) Radio equipment - OFF. FUEL INTCON switch -

NORM.

FUEL TRANS switches - OFF.

BOOST PUMP switches - OFF. FUEL XFEED switch - NORM. Rev. 19

2-7

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O ENGINE 1 and ENGINE 2 FUEL switches - OFF.

VENT BLOWER switch EMERG

PART SEP switches -

DISARM.

NORM.

ENGINE 1 and ENGINE 2 GOV switches - AUTO. HYDR SYS NO.1 and NO.2 switches ON.

-

STEP switch (if installed) - As desired.

LT switch

(if

OFF.

installed)

STBY ATT switch (if installed) - TEST; check standby attitude indicator light and OFF flag retracts illuminates momentarily, then switch OFF. WSHLD HEAT switches OFF.

(if installed) -

Overhead circuit breakers -

ROTOR RPM audio loaded to AUDIO.

switch

FORCE TRIM switch Instruments -

-

Spring

ON, cover down.

All LT rheostats -

STATIC SOURCE switch (if installed) PRI.

OFF.

OFF.

ANTI COLL light -

ON.

EMERG LOAD switch -

NORMAL.

APPROACH PLATE AND MAP LIGHT knob(s) - OFF.

BUS switch NON-ESNTL loaded to NORMAL.

AUX SYS PITOT and STATIC switches (if installed) - NORM.

INV 1 and 2 switches -

Altimeter(s) Clock -

In.

OFF.

UTILITY LIGHT switch POSITION light -

Static check.

-

GEN 1 and 2 switches -

-

Spring

OFF.

OFF.

- Set.

IF EXTERNAL POWER IS USED CONNECT (1000 AMPS MAXIMUM). CHECK 27 ± 1 VDC; ADJUST POWER SOURCE IF REQUIRED.

Set and running.

FIRE EXT switch -

OFF.

FIRE PULL handles -

In (forward).

AFT DOME LIGHT rheostat and switch OFF. PITOT STATIC HEATERS switch WIPER switches -

OFF.

OFF.

CARGO RELEASE switch (if installed) OFF. HEATER switch -

OFF.

AFT OUTLET switch -

OFF.

BATTERY BUS 1 and BUS 2 switches light ON; check BATTERY caution illuminates.

NOTE

Test operate all lights when night flights are planned or anticipated. with light tests Accomplish or power connected external during engine runup.

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O NOTE If Anticollision light is desired, NON-ESNTL BUS -

MANUAL.

NOTE

Rotor brake shall be off at all times when the engines are running.

FIRE 1 and 2 warning lights test button Press to test.

BAGGAGE FIRE warning light test button - Press to test (verify light flashes).

OVER TORQ caution light - Press. Check light illuminates and MAST TORQUEindicator reads 105 ± 1%. CAUTION

IF MAST TORQUE INDICATOR INDICATESAN ERROR GREATER THAN ± 1% FROM THE 105% POSITION, THE MAST TORQUE SYSTEM IS UNRELIABLE. MAINTENANCE ACTION IS REQUIRED.

CYC CTR caution lights - Press to test.

Caution panel TEST switch - PNL (All segments extinguish except CAUTION PANEL). Caution panel TEST switch segments illuminate).

LT (All

Caution panel RESET button - Press (MASTERCAUTIONlight extinguishes).

FUEL SYS test switch - FWD TANK, then MID TANK; note digital and needle indications. FUEL SYS DIGITS TEST button (digital display reads 888).

ON.

2-5. ENGINESTART

ROTOR BRAKE lights - Test. Pull brake ON and check that both lights illuminate; return to OFF and check lights extinguish.

-


Press

NOTE

If the helicopter has been cold soaked in ambient temperatures of -18°C (0°F) or less, both throttles will be difficult to move and follow through coupling may be increased.

Throttles - Rotate engine 1 throttle full open, then back against idle stop. Actuate ENG 1 IDLE STOP release, roll engine 1 throttle to full closed, then apply friction as desired. Repeat procedure using engine 2 throttle and ENG 2 IDLE STOP release. NOTE

When either IDLE STOP release is activated, the appropriate idle stop plunger will not release if pressure is applied toward the closed position of the throttle.

Moderate frictions should be applied to overcome follow-through coupling between throttles. RPM INCR/DECRswitch seconds.

DECR for 8

2-5-A. ENGINE 1 START NOTE

Either engine may be started first. However, the following procedure is provided for starting engine 1 first. At ambient temperatures 5°C (41°F) and below, refer to paragraph 2-5-F.

Engine 1 FUEL TRANS switch - ON, check NO. 1 FUEL TRANS caution light extinguished. Rev.17

2-9

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S I N G O N I R W O E I F V NLY O Engine 1 BOOST PUMP switch - ON, check NO. 1 FUEL BOOST light extinguished. Engine 1 FUEL switch - ON. (FUEL VALVE caution light will illuminate momentarily.) Engine 1 FUEL PRESS indicator Check. Rotor -

-

On side slopes greater than five degrees, disregard CYC CTR caution lights and position cyclic as desired.

Cyclic - Position as necessary extinguish CYC CTR caution lights.

PROLONGED EXPOSURE TO AMBIENT TEMPERATURES OF 0°C (32°F) OR LESS MAY FREEZE MOISTUREIN THE ENGINE FUEL CONTROL SYSTEM. MONITOR ENG RPM (N) DURING COLD FOR WEATHER STARTING OVERSPEED.IF AN OVERSPEED APPEARS IMMINENT, ABORT START AND CLOSE THROTTLE TO OFF POSITION.

Engine 1 ENGINE OIL pressure Indicating.

NOTE

START switch - Centered (off) at 55% GAS PROD RPM(N,).

-

Engine 1 throttle - Open to idle at 12% GAS PROD RPM(N,) minimum. Engine 1 ITT - Monitor to avoid a hot start. Maximum ITT during start is 1090°C, not to exceed two seconds above 960°C. If ITT continues to rise, abort start by activating idle stop release and rolling throttle to fully closed. Starter should remain engaged until ITT decreases. Do not attempt restart until corrective maintenance has been accomplished. NOTE

If engine fails to start, refer to paragraph 2-5-D procedures.

Collective pitch position.

to

CYC CTR caution lights are inhibited between 95 and 105% ROTORRPM.

START switch - ENG 1 position. Observe starter limitations.

Rev.17

IF STICK CENTERINGINDICATOR SYSTEM IS INOPERATIVE, GROUND OPERATIONSHALL BE CONDUCTED AT 97% ROTOR RPM OR ABOVE. NOTE

Clear.

CAUTION

2-10

CAUTION

Ensure in full down

GAS PROD - Check 61 + 1% when throttle is on idle stop. NOTE Do not increase ROTOR above XMSN oil 85% RPM until temperature is above 15°C.

ENGINE, XMSN, and GEAR BOX OIL PRESSURES- Check. Engine 1 PART SEP OFF caution light Check extinguished.

CAUTION

DURING RPM INCREASE, ANY ABNORMAL INCREASE IN ONEMAY PER-REV VIBRATION INDICATE ONE OR MORE MAIN DROOP RESTRAINERS FAILED TO DISENGAGE FROM STATIC POSITION. VERIFY PROPER OPERATIONPRIORTO FLIGHT.

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Engine 1 throttle -

increase to 77 - 85%

ENG RPM(Ni). Friction as desired. NOTE

Minimize blade flapping by maintaining highest rotor RPM (NR)within allowable range.

Rotor RPM (NR) -

desired.

Minimum 77 - 85%, as

If external power is used, proceed to engine 2 start. If battery was used, proceed as follows: Check 71% minimum.

GEN 1 switch -

ON.

Engine 2 ENGINE OIL pressure Indicating.

-

Engine 2 throttle - Open to idle at 12% GAS PROD RPM(N,). Monitor. Observe ITT

START switch - Centered (off) at 55% GAS PROD RPM(N,). GAS PROD - Check 61+1% RPM (N,) when engine 2 throttle is on Idle stop. CAUTION

NOTE

Before attempting generator assisted start on second engine, it is recommended that the battery be charged until the ammeter load drops below 150 amps.

AMPS 1 indicator 150 amps.

Check.

START switch - ENG 2 position. Observe starter limitations.

Engine 2 ITT limitations.

NOTE

N, RPM-

Engine 2 FUEL PRESS-

Check at or below

2-5-B. ENGINE 2 START NOTE

At ambient temperatures of 5°C (41°F) and below, refer to paragraph 2-5-F.

Engine 2 FUEL TRANS switch - ON. Check NO. 2 FUEL TRANS caution light extinguished. Engine 2 BOOST PUMP switch - ON. Check No. 2 FUEL BOOST light out (FUEL XFEED caution light will illuminate momentarily.) Engine 2 FUEL switch - ON (FUEL VALVE caution light will illuminate momentarily).

ENSURE SECOND ENGINE ENGAGES AS THROTTLE IS INCREASED. A NONENGAGED ENGINE INDICATES 10 TO 15% HIGHER N,, RPM THAN THE ENGAGED ENGINE AND NEAR ZERO TORQUE. IF A NONENGAGEMENT OCCURS, CLOSE THROTTLE OF THE NONENGAGED ENGINE. WHEN THE NONENGAGEDENGINE HAS STOPPED, SHUT DOWN THE ENGAGEDENGINE. IF A SUDDEN (HARD) ENGAGEMENT OCCURS, SHUT DOWN BOTH ENGINES. MAINTENANCEIS REQUIRED.

Engine 2 throttle - Increase slowly to match Engine 1 N,, rpm. Monitor tachometer and torquemeter to verify engagement of second engine. Engine 2 oil pressure -

Check.

ENG 2 PART SEP OFF caution light Check extinguished.

2-5-C. POST START Rev. 17

2-11

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Radios -

NR-

77 - 85% or above. NOTE

If external power was used Disconnect. GEN 1 switch - ON.

GEN 2 switch - ON. (BAT BUS 1 will switch off automatically.)

ON as required.

ELT (if installed) inadvertent transmission.

Check

for

2-5-D. ENGINEFAILS TO START

When the engine fails to start within 15 seconds after the throttle has been opened to idle, the following action is recommended:

CAUTION

IDLE STOP release - Actuate.

ONLY BATTERY BUS 2 SWITCH SHOULD REMAIN ON WITH BOTH GENERATORSOPERATING. NOTE

any rpm increase, During abnormal increase in 1 per rev vibration may indicate one or more main rotor droop restraints failed to disengage from static position. Verify proper operation prior to operating.

Caution lights (except AFCS).

Check all extinguished

ENGINE, XMSN, and GEAR BOX OIL temperatures and pressures - Within limits. AMPS 1 and 2 -

Within limits.

NOTE

AMPS 2 will indicate a higher load than AMPS 1 until battery is fully charged.

2-12

Rev.17

Throttle Starter -

Fully close.

Disengage.

BOOST PUMP switch - OFF. FUEL switch -

OFF.

After GAS PROD RPM (N1) has decreased to zero, allow 30 seconds for fuel to drain from engine.

Conduct a DRY MOTORINGRUN before attempting another start.

2-5-E. DRY MOTORINGRUN

The following procedure is used to clear an engine whenever it is deemed necessary to remove internally trapped fuel and vapor. Throttle -

Fully closed.

BOOST PUMPswitch -

ON.


FAA APPROVED

FUEL switch - ON. IGN circuit breaker-

Pull out.

Starter - Engage for 15 seconds, then disengage.

At ambient temperature 5°C (41 F) and below, incorporate warmup procedures as follows:

BOOST PUMP switch - OFF. FUEL OFF. FUEL switch switch -- OFF. IGN circuit breaker -

ENG 1/2 throttle -

Push in.

Allow the required cooling period for the starter before proceeding. Follow normal start sequence as described on preceding pages.

2-5-F.

COLD WEATHER START

Open slowly to

maintain 77 - 85% ENG RPM as desired (minimum 3 minutes).

After warm-up:

Throttle - Reduce to idle. GOV switch - Return to AUTO.

This procedure is applicable to either engine.

CAUTION

DO NOT ALLOW GAS PRODUCER TO DECREASE BELOW 50% RPM (N1). NOTE

ambient cold extremely During temperatures, idle rpm will be high and the ENGINE, XMSN, and GEAR BOX OIL pressure may exceed limits for up to two minutes after starting. Warm up shall be conducted at 77 - 85% NR at

flat pitch.

Continue start sequence per paragraph 2-5-A or 2-5-B.

Rev. 20

2-12A / 2-12B


BHT-412-FM-4

2-6. SYSTEMS CHECK

2-6-B. FORCE TRIM CHECK

2-6-A.

2-6-A. STICK STICK CENTERING CENTERING INDICATOR CHECK

Flight controls lock lock removed. removed.

Cyclic and pedals - Move slightly each direction to check force gradients.

CAUTION

CAUTION

Yaw trim release switch (if installed) Press forward. Check yaw trim only releases, reengages with switch released.

DURING EXTREME COLD AMBIENT TEMPERATURES LIMIT CYCLIC MOVEMENTS UNTIL XMSN OIL TEMPERATURE REACHES 15°C.

CAUTION

Friction off, collective

Cyclic FORCE TRIM release button Depress. Check trim releases with button depressed, reengages with button released.

adisengages

FORCE TRIM switch - OFF. Check trim and FT OFF caution light illuminates. FORCE TRIM switch - ON, cover down.

DO NOT DISPLACE CYCLIC MORE THAN 1.5 INCHES FROM CENTER TO CHECK THE SYSTEM. IF CYC CTR CAUTION LIGHTS DO NOT ILLUMINATE WITHIN THE 1.5 INCH DISPLACEMENT, THE SYSTEM IS INOPERATIVE. DO NOT DISPLACE CYCLIC BEYOND POINT AT WHICH CYC CTR CAUTION LIGHT ILLUMINATES. NOTE

2-6-C. PRELIMINARY HYDRAULIC CHECK Throttles -

Set to idle. NOTE

Uncommanded control movement or motoring with either hydraulic system off may indicate hydraulic system malfunction. NO. 1 HYD SYS switch OFF, then ON. NO. 2 HYD SYS switch -

OFF, then ON.

CYC CTR caution lights are inhibited between 95 and 105% ROTOR RPM.

.

Cyclic - Displace approximately 1.25 Inch (31.7 mm) forward, aft, left and right. Check CYC CTR caution light illuminates each time when displaced and extinguishes when centered.

Rev. 17

2-13

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAAAPPROVED

BHT-412-FM-4

2-6-D.

ENGINE FUEL CONTROL

CHECK

Throttles (both) - Idle.

CAUTION

DO NOT ALLOW GAS PROD TO DECREASE BELOW 50% RPM (N.).

NOTE In the vicinity of 8000 feet pressure altitude, GAS PROD RPM (N.) may not change

significantly when manual fuel control is selected.

GOV switch (ENGINE 1 or 2) - MANUAL observe a change in GAS PROD RPM (N|) and GOV MANUAL caution light illuminates.

Open respective throttle carefully to ensure GAS PROD RPM (N|) responds upward, then return to idle position.

Return GOV switch to AUTO. Check for a return to original GAS PROD RPM (N|) and GOV MANUAL caution light extinguishes. Check second governor in like manner.

Throttles (both) - Increase slowly above 85% ROTOR RPM.

2-6-E.

GOVERNOR CHECK CAUTION

IF HELICOPTER IS SITTING ON ICE OR OTHER SLIPPERY OR LOOSE SURFACE, ADVANCE THROTTLES SLOWLY TO PREVENT ROTATION OF HELICOPTER.

Engine 1 throttle - Fully open. ENG - Stabilized at 95 ± 1% RPM (N||). 2-14

Rev. 22 11 JUN 2001

Engine 2 throttle - Full open. Check No. 1 engine Increases 2% ENG RPM (Ni,) and both engines stabilize at 97 ± 1% ENG RPM (Nil). (NOTE Helicopters rigged for Category A operations will obtain 103% ENG RPM (Nn). Refer to Flight Manual Supplement BHT412-FMS-62.3 & 62.4 for Airspeed Limitations.

RPM INCR/DECR switch - Full INCR. Check ENG does not exceed 101.5% RPM (Nn). Set at 100% ENG RPM (Nn).

2-6-F. FUEL CROSSFEED AND INTERCONNECT VALVE CHECK FUEL XFEEDNTCON test switch BUS 1 and hold.

TEST

NOTE After turning either boost pump off, FUEL BOOST caution light should illuminate on failed side only.

ENGINE 1 BOOST PUMP switch - OFF. Check Engine I FUEL PRESSURE decreases, then returns to normal. (This indicates crossfeed valve has been opened by Bus No. 1 power and check valve is functioning properly). Return switch to ON. FUEL INTCON switch - OPEN. Check FUEL INTCON caution light Illuminates then extinguishes. (This Indicates interconnect valve has been opened by BUS NO. 1 power and that the valve is functioning properly). FUEL INTCON switch - OVRD CLOSE. Check FUEL INTCON caution light illuminates then extinguishes. FUEL XFEED/INTCON test switch - TEST BUS 2 and hold.

ENGINE switch - OFF. PRESS decrease ENGINE 22 BOOST BOOST PUMP Check engine 2 FUEL PRESS decreases, then returns to normal. Return switch to ON.

BHT-412-FM-4


FAA APPROVED FUEL INTCON switch - OPEN. Check FUEL INTCON caution light illuminates then extinguishes. (This indicates that the interconnect valve has been opened by BUS NO. 2 power and that the valve is functioning properly). FUEL INTCON switch - NORM. Check FUEL INTCON caution light illuminates then extinguishes. FUEL XFEED/INTCON test switch NORM. FUEL XFEED switch -

OVRDCLOSE.

EMERGLOAD switch -


BHT-412-FM-4 NORMAL ON

INV 1 switch - OFF; check INVERTER1 caution light illuminates. Check No. 1 and No. 2 AC VOLTS for indication that inverter 2 has assumed all ac loads. Return INV 1 switch to ON. INV 2 switch - OFF; check INVERTER2 caution light illuminates. Check No. 1 and No. 2 AC VOLTS for indication that inverter 1 has assumed all ac loads. Return INV 2 switch to ON.

ENGINE 1 BOOST PUMP switch - OFF. Check fuel pressure drops to zero on affected system. Return switch to ON. Repeat procedure for ENGINE 2 BOOST PUMPswitch.

STBY ATT switch (if installed) - ON.

FUELXFEED switch -

glow.

NORM.

2-6-G. ELECTRICAL SYSTEMS CHECK DC VOLTS-

Check 27 ± 1 volts.

AC VOLTS-

Check 104 to 122 volts

AMPS 1 and 2 - Check within limits. GEN 1 and 2 switches -

OFF.

EMERG LOAD switch - EMERG LOAD. Check that the following items remain operational:

EMERGLT switch (if installed) - TEST; check all emergency lights illuminate. Switch to ARM; check lights dim to faint

2-6-H. HYDRAULIC SYSTEMS CHECK NOTE

The Hydraulic Systems Check is to determine proper operation of the hydraulic actuators for each flight control system. If abnormal forces, unequal forces, control binding or motoring are encountered, it may be an indication of a malfunctioning flight control actuator.

AP 2.

FORCE TRIM switch -

OFF.

VHF #1.

Collective removed.

down,

-

Full

friction

Pilot map light

Rotor -

Set to 100% RPM.

Cyclic -

Centered, friction removed.

Pilots ICS

STBY ATT indicator

Essential engine instruments (Refer to BHT-412-MD-4 for a complete list of EMER BUS components).

HYDR SYS NO. 1 switch - OFF. Check NO. 1 HYDRAULIC caution light and MASTER CAUTION light illuminate and system 1 pressure drops to zero. Rev.9

2-15

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Cyclic - Check normal operation by moving cyclic forward, aft, left, and right approximately one inch. Center cyclic.

Collective - Check for normal operation by increasing collective control 1 to 2 inches. Repeat 2 to 3 times as required. Return to full down position. Pedals - Displace slightly left and right. Note an increase in force required to move pedal in each direction.

An electrical interlock prevents both hydraulic systems from being switched off at the same time. If one system is off and the second system is switched off, the second system will remain on.

HYDR SYS NO. 2 switch - OFF. Check hydraulic system 2 remains operational and system 1 remains off.

HYDR SYS NO. 1 switch - ON. Check light caution NO. 1 HYDRAULIC and system 1 regains extinguishes NO. 2 Check normal pressure. HYDRAULIC caution light illuminates and system 2 pressure drops to zero. Cyclic - Check normal operation by moving cyclic forward, aft, left, and right approximately one inch. Center cyclic.

Collective - Check for normal operation by increasing collective control 1 to 2 inches. Repeat 2 to 3 times as required. Return to full down position. Pedals - Displace slightly left and right. Note the pedals are now boosted.

HYDR SYS NO. 2 switch - ON. Check light caution NO. 2 HYDRAULIC extinguishes, system 2 pressure returns system 1 to normal, and hydraulic remains operational.

2-16

Rev. 19

WARNING

BOTH HYDRAULIC SYSTEMS SHALL BE OPERATIONAL PRIOR TO TAKEOFF. NOTE

System 1 will normally operate 10 to 20 °C cooler than system 2.

NOTE

Cyclic and collective friction desired.

FORCE TRIM switch - ON. Check FT OFF caution light extinguishes.

Set as

2-6-J.

AFCS CHECK

NOTE

Verification of AFCS actuator centering is necessary. Failure of the actuators to center could result in reduced control margins (pitch and roll) and abnormal control positions (pitch, roll, and yaw). NOTE

For S/N 36087 - 36247, if fast slaving is desired, center ADI roll trim knob, then push and hold VG FAST ERECT button until attitude indicator displays zero degrees bank angle. Use of VG FAST ERECT button will disengage the respective autopilot.

Pilot and copilot attitude indicators Erect and set as necessary.

WARNING

IF AFCS IS LEFT ENGAGED IN ATT MODE DURING GROUND OPERATION, IT CAN DRIVE THE CYCLIC STICK TO A CONTROL STOP.

I


FAA APPROVED

Verify AP 1 and AP 2 are disengaged.

completion of AFCS CHECK.

Cyclic force trim switch - Check ON.

Press TEST switch to exit AFCS test.

For helicopters not equipped with Flight Director Kit: Press and hold the CPL switch on autopilot control panel. While holding CPL switch, press the TEST switch. Release both

2-6-K.

switches - verify TESTON.

Roll adjustment indicies aligned, press and

Confirm AFCS TEST flashing, ADI Pitch and hold AHRS control TEST switches.

NOTE

will light AUTO TRIM illuminated during AFCS test.

AHRS CHECK (S/N 36248 AND SUBSEQUENT)

Panel No. 1 and No. 2

remain

NOTE

Test may be performed in MAG or DG mode.If TEST accomplished in DG mode, heading must be manually aligned.

AFCS TEST switch -

ON for level 1 test.

Verify following responses for each system:

Verify AL-300 Air Data Display (ADD) reads LE 1.

HSI

Verify AP 1 light flashes, then press.

Heading

Rotates to 015°

Heading OFF Flag

In view

NOTE

Pilot shall closely monitor controls during test, taking care not to restrict control movements. To prevent nuisance failures, hands must remain

ADI

Pitch

5° nose up

Roll

45° right

Cyclic-- Verify movement aft right approximately one inch, then return to neutral.

ATT Flag

In view

Verify AP 2 light flashes, then press.

RT Flag

In view

Cyclic - Verify movement aft right approximately one inch, then return to neutral.

AHRS CONTROL PANEL

Upon completion of test, verify AP 1 and AP 2 lights extinguish and AL-300 reads END 1.

* / + Indicator

off controls.

If error codes are present, press SBY (or cyclic NAV STBY switch) once to view each code until END 1 is displayed (refer to table 3-3 for a list of error codes). (AHRS equipped helicopters) - Proceed to paragraph 2-6-K AHRS CHECK prior to

Needle Deflection - *

Press AFCS TEST switch to exit test mode. Rev. 20

2-17

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4 2-6-L.

FAA APPROVED

Within operating ranges.

CABIN HEATER CHECK

GAS PROD Check 75% RPM (N1) minimum (both engines).

Caution and Extinguished.

Thermostat knob - Full COLD.

Flight instruments

warning -

lights

-

Check operation

and set. When flying IFR check as follows:

CAUTION

IVSI - Needle near zero.

HEATER SWITCH SHALL BE TURNEDOFF WHEN HEATED AIRFLOW DOES NOT SHUT OFF AFTER THERMOSTAT IS TURNED TO FULL COLD, HEATER AIR LINE LIGHT ILLUMINATES, OR CABIN HTR CIRCUIT BREAKER TRIPS.

Altimeter - Set and check. HSI - Check.

ADI - Check.

Radio Altimeter - Zero altitude, OFF flag retracted, DH set as desired.

CAUTION

WARNING

DO NOT OPERATE HEATER ABOVE 21 °C OAT.

VENT BLOWER switch -

ON.

Thermostat setting observe heated airflow. DEFOG lever -

Increase

IF AFCS IS LEFT ENGAGED IN ATT

MODE DURING GROUND OPERATION, IT CAN DRIVE THE CYCLIC STICK TO A CONTROL STOP.

HEATER switch - ON.

and

ON. Check airflow is

diverted from pedestal outlets to windshield nozzles. Return lever to OFF.

AFT OUTLET switch - ON. Check airflow distributed equally between pedestal outlets and aft outlets. Return switch to OFF.

AFCS - AP 1 and AP 2 engaged and select ATT or SAS mode as desired (ATT

mode shall be used during IFR flight;

SAS mode recommended for ground operations, hover,andtakeoff). TRIM/FD 1 or 2 - As desired. desired.

TRIM/FD 1 or 2-As FORCE TRIM Switch - ON in ATT mode, as desired in SAS mode. Auxiliary navigation Panel - As desired.

NOTE

Heater operation affects performance. Refer to Hover Ceiling and Rate of Climb charts for HEATER ON in Section 4. HEATER switch - As desired. VENT BLOWER switch -

2-7.

BEFORE Engine,

As desired.

TAKEOFF

gearbox,

Rev. 19

Throttles -

Full open. Adjust frictions.

ENG - 100% RPM (N11)for both engines. transmission,

hydraulic, and electrical instruments 2-18

MODERATE FRICTION SHALL BE APPLIED TO OVERCOME FOLLOWBETWEEN THROUGH COUPLING THROTTLES.

POSITION lights -

As required.

FAA APPROVED

BHT-412-FM-4

ANTI COLL light — Check ON.

NR — 77 - 85% or above, as desired.

PITOT-STATIC HEATERS switch — ON. Check ammeter for load indication. Leave ON in visible moisture when temperature is below 4.4°C (40°F) OFF if not required.

Cyclic — Position as necessary to extinguish CYC CTR caution lights.

Radio(s) — Check functioning.

On side slopes greater than five degrees, maintain 100% ROTOR RPM. CYC CTR caution lights are inhibited.

NOTE

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O EMERGENCY COMM panel — (If installed) Check for single pilot operations. Cyclic control — Centered or slightly into the wind.

COCKPIT VOICE RECORDER TEST switch (if installed) — Press and hold for three seconds. Verify meter indicates GOOD.

2-7-C.

OPERATION IN OR NEAR MAGNETIC DISTURBANCES (S/N 36248 AND SUB)

NOTE

For other Cockpit Voice Recorder installations, refer to the applicable Flight Manual Supplement.

STEP switch (if installed) — As desired. Passenger seat belts — Fastened. All doors — Secured.

2-7-A.

POWER ASSURANCE CHECK

Power assurance check should be performed daily. (Refer to Section 4.)

2-7-B.

PROLONGED GROUND OPERATION NOTE

For prolonged ground operation, AFCS shall not be operated in ATT mode.

CAUTION

TO AVOID HEADING ERROR, ACCOMPLISH INFLIGHT AHRS HEADING SYNCHRONIZATION ONLY IN LEVEL FLIGHT AND, IF FLIGHT DIRECTOR INSTALLED, UNCOUPLED (STBY).

When operation AHRS equipped helicopters in an area of suspected adverse magnetic disturbance (e.g., an off-shore oil platform), a significant heading error can occur during landing and takeoff. Prior to landing or before takeoff, switch AHRS from MAG mode to DG mode and, if necessary, manually align heading with visual reference. When clear of magnetic disturbance switch back to MAG mode to initiate automatic heading synchronization.

2-8.

TAKEOFF

CAUTION

CAUTION

MINIMUM ROTOR — 97% RPM FOR GROUND OPERATION WITH STICK CENTERING INDICATOR SYSTEM INOPERATIVE.

DURING LIFTOFF TO HOVER, ANY ABNORMAL INCREASE IN ONE PER REV VIBRATION MAY INDICATE ONE OR MORE MAIN ROTOR DROOP RESTRAINERS FAILED TO DISENGAGE FROM STATIC POSITION. VERIFY PROPER OPERATION PRIOR TO FLIGHT.

NOTE For ground operation, maintain rotor RPM (N R ) within allowable range. Higher minimum rotor RPM reduces blade flapping.

23 MAY 2002

Rev. 23

2-19


FAA APPROVED

NOTE

When AFCS is in ATT mode, the FORCE TRIM release button and the CLTV/YAW TRIM REL switch should be depressed before liftoff (to trim actuators to center positions) and should be held until desired climbout attitude is attained. AP 1 and AP 2 - As required TRIM/FD 1 or 2 - As desired. Area-

AVOID area of the Height-Velocity diagram. Torque shall not exceed 15% above IGE hover power while accelerating to Takeoff Climbout Safety Speed. (Refer to Section 4.) Cyclic control - Apply forward cyclic to

Collective pitch - Adjust as desired after reaching VTOCS (45 KIAS). Airspeed - Within limits (60 KIAS minimum for IFR).

2-9.

Clear.

IN-FLIGHT OPERATIONS

(AHRS equipped helicopters - S/N 36248 and subsequent). CAUTION

NOTE HSI HEADING PRIOR TO LIFTOFF. TO ENSURE NO MAGNETIC DISTURBANCE EXISTS, CROSSCHECK HEADING IN HOVER. IF DISCREPANCY EXISTS, OPERATE IN DG MODE (REFER TO OPERATIONS IN OR NEAR MAGNETIC DISTURBANCES). DISTURBANCES). NOTE

As collective is increased, it may be necessary to rematch engine torques prior to reaching hover.

RPM (trim) switch - Adjust to obtain matching torque, N1,or ITT, as required, and 100% NR.

Hover power - Check torque required to hover at four feet skid height. NOTE

not are takeoffs Downwind recommended since the published takeoff distance performance will not be achieved.

During takeoff, pitch attitude must be adjusted commensurate with power application to prevent entering the 2-20

Rev. 20

CAUTION

WHEN OPERATING NEAR THE MAXIMUM MAST TORQUE LIMIT, INADVERTANT OVERTORQUE MAY OCCUR DURING MANEUVERING FLIGHT CONDITIONS INVOLVING TURNS AND/OR NOSE DOWN ATTITUDE CHANGES. DECREASE POWER TO 90% MAST TORQUE PRIOR TO MANEUVERING HELICOPTER. ENG - 97 to 100% RPM(N11). Airspeed - Within limits for flight altitude.

Engine, gearbox, and transmission Instruments - Within limits.

NOTE Maximum pitch attitude capability of standby attitude indicator is ±60

degrees.

Refer to applicable operating rules for high altitude oxygen requirements. (AHRS equipped helicopters - S/N 36248 and subsequent).

CAUTION

WHEN AHRS IS SWITCHED FROM DG IT MAG MODE


BHT-412-FM-4

INSTANTANEOUSLY SYNCHRONIZE TO DETECTED MAGNETIC HEADING. MAKE SWITCH SELECTION ONLY IN LEVEL FLIGHT TO AVOID INVALID HEADING SYNCHRONIZATION DUE TO FLUX VALVE SENSING ERRORS DURING TURNS.

2-9-A.

MANEUVERING SAS MODE

desired (minimum friction for ATT mode operation). AFCS desired.

WITH AFCS IN

MANEUVERING WITH AFCS IN ATT MODE

For momentary yaw changes, due to power change, press forward on YAW TRIM REL switch (if installed), which releases pedal force trim only. Adjust yaw with pedals, then release switch. For momentary attitude changes, manual cyclic movement may be used. However, AFCS actuators may be saturated to limit authority when cyclic is moved manually.

Adjust

CAUTION

RUN ON LANDINGS MAY RESULT IN ROLL OSCILLATIONS WHILE ON THE GROUND. IF THIS OCCURS, LOWERING COLLECTIVE FULLY DOWN OR DISENGAGING AP 1 AND AP 2 WILL STOP THE OSCILLATIONS. AFTER LANDING Collective pitch -

Full down.

FORCE TRIM release button - Depress and check for actuator centering.

DESCENT AND LANDING -

STEP SWITCH (if installed) - As desired.

FORCE TRIM switch - ON.

In-flight use of VG FAST ERECT button (S/N 36087 - 36247) will disengage respective autopilot.

controls

Flight path - Stay clear of AVOID area of Height-velocity diagram (Refer to Section 1). For landing distance Information in the event of engine failure during approach (refer to Section 4).

Pedals - Centered.

NOTE

Flight

Full open.

ENG - 100% RPM (Nl).

Depress cyclic FORCE TRIM release button and maneuver as desired. Release button when desired attitude is reached. Autopilot will hold attitude until retrimmed to new attitude. Attitude may also be adjusted with cyclic ATTD TRIM switch.

2-10.

FORCE TRIM switch - ON in ATT mode, as desired In SAS mode. Throttles -

Use normal pilot control techniques.

2-9-B.

Engage ATT or SAS mode as

friction

AFCS - SAS mode.

as

Rev. 19

2-20A/2-20B


BHT-412-FM-4

2-11. ENGINESHUTDOWN

AP 1 and AP 2 - Disengage, check autopilot lights extinguish, AFCS and MASTER CAUTIONlights illuminate. Cyclic - Friction as desired. Maintain cyclic stick as near center as possible at all rotor speeds. NOTE For ground operation, maintain rotor RPM (NR) within allowable range. Higher minimum rotor RPM reduces blade flapping.

Throttles -

desired.


OFF.

OFF.

Engine 1 and 2 FUEL switches -

Reduce to 77 - 85% NR, as

ENG 2

Engine 2 throttle - Fully closed. Check ITT and GAS PRODRPM (N,) decreasing. GEN 1 and 2 switches -

OFF.

Engine 1 and 2 BOOST PUMP switches -

OFF.

Engine 1 and 2 FUELTRANS switches OFF.

CAUTION

Radios -

IF STICK CENTERINGINDICATOR SYSTEMIS INOPERATIVE,NR FOR GROUND OPERATIONSHALL BE 97% OR ABOVE.

NOTE On side slopes greater than five degrees, disregard CYC CTR caution lights and position cyclic as required. Check

STBY ATTD switch (if installed) -

for

OFF.

EMERG LT switch DISARM.

(if installed)

Engine instruments -

Within limits.

DO NOT USE COLLECTIVE TO SLOW ROTOR RPM. USE OF COLLECTIVE TO SLOW ROTOR RPM CAN CAUSE EXCESSIVE FLAPPINGAND / OR CONING. CAUTION

Stabilize for one minute.

ELT (if installed) inadvertent transmission.

OFF.

WARNING

Stick centering check - Complete. Center cyclic and friction as necessary to extinguish CYC CTR caution lights.

ITT -

IDLE STOP release switch position.

-

AVOID RAPID ENGAGEMENTOF ROTOR BRAKE IF HELICOPTERIS ON ICE OR OTHER SLIPPERY OR LOOSE SURFACE TO PREVENT ROTATIONOF HELICOPTER.

Rotor brake - As desired. Apply at or below 40% ROTOR RPM. Return to stowed position after main rotor stops. Pilot - Remain at flight controls until rotor has come to a complete stop.

ENG 1

Lighting and miscellaneous switches OFF.

Engine 1 throttle - Fully closed. Check ITT and GAS PROD RPM (N,) decreasing.

BATTERY BUS 1 and BUS 2 switches OFF

IDLE STOP release switch position.

BATTERY BUS 1 Switch -

ON.

Collective desired.

down lock -

Secured as

Rev. 15

2-21


FAA APPROVED

droop restraint arms are engaged in the lower detent of the cam window.

2-12. POSTFLIGHTCHECK If conditions following:

require,

perform

the

NOTE

Refer to BHT-412-MD-4 additional information.

for

Check general condition of droop restraint system and verify that the

Rev.13

Install main rotor blade tiedown socks on blades and secure to mooring points. Install tail rotor tiedown secure to vertical fin.

strap and

Install exhaust covers, engine inlet protective plugs, and pitot tube covers.


FAA APPROVED

BHT-412EP-FM-2-1

Figure 2-1.

Preflight check sequence

2-23/2-24

T Section 3 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FM-4

TABLE OF CONTENTS

Subject

INTRODUCTION ......................................................................... DEFINITIONS .............................................................................. ENGINE ....................................................................................... SINGLE ENGINE FAILURE .................................................... ENGINE RESTART IN FLIGHT............................................... DUAL ENGINE FAILURE........................................................ ENGINE UNDERSPEED ......................................................... ENGINE OVERSPEED- FUEL CONTROL / GOVERNOR FAILURE ................................................................................. ENGINE OVERSPEED- DRIVESHAFT FAILURE............... ENGINE COMPRESSORSTALL............................................ ENGINE HOT START/SHUTDOWN ....................................... FIRE............................................................................................. ENGINE FIRE .......................................................................... CABIN SMOKE OR FUMES ................................................... BAGGAGE COMPARTMENT FIRE........................................ TAIL ROTOR............................................................................... COMPLETE LOSS OF TAIL ROTOR THRUST...................... TAIL ROTOR FIXED PITCH FAILURES................................ LOSS OF PITCHCHANGE CONTROL LINKAGE ................. HYDRAULIC SYSTEM................................................................ ELECTRICAL SYSTEM .............................................................. DC POWER FAILURE............................................................. AC POWER FAILURE............................................................. FUEL SYSTEM............................................................................ FUEL BOOST PUMP FAILURE................................. FUEL FILTER PARTIALLY BLOCKED .................................. FUEL QUANTITY INDICATIONS MALFUNCTION................ AUTOMATIC FLIGHT CONTROLS SYSTEM............................ AFCS FAILS TO ENGAGE OR DISENGAGE ........................ AFCS FAILS TO HOLD ATTITUDE........................................ AFCS HARDOVER OR ABNORMAL CONTROL DISTURBANCE....................................................................... AUTOTRIM FAILURE ............................................................. PITCH OR ROLL FAILURE .................................................... AUTOTRIM RUNAWAY .......................................................... AHRS - ALIGNMENT IN FLIGHT (S/N 36248 AND SUB) ... COMMUNICATION SYSTEM...................................................... INTERCOM FAILURE .............................................................

Paragraph

Page Number

3-1 .................. 3-2.................. 3-3 .................. 3-3-A .............. 3-3-B .............. 3-3-C.............. 3-3-D..............

3-3 3-3 3-3 3-3 3-5 3-6 3-7

3-3-E .............. 3-3-F .............. 3-3-G.............. 3-3-H .............. 3-4 .................. 3-4-A .............. 3-4-B .............. 3-4-C .............. 3-5.................. 3-5-A .............. 3-5-B .............. 3-5-C .............. 3-6.................. 3-7.................. 3-7-A .............. 3-7-B .............. 3-8 .................. 3-8-A .............. 3-8-B .............. 3-8-C .............. 3-9.................. 3-9-A .............. 3-9-B ..............

3-8 3-9 3-9 3-10 3-10 3-10 3-11 3-12 3-12 3-12 3-14 3-15 3-16 3-16 3-16 3-17 3-18 3-18 3-18 3-18 3-18 3-19 3-20

3-9-C .............. 3-9-D .............. 3-9-E .............. 3-9-F .............. 3-9-G.............. 3-10................ 3-10-A ............

3-20 3-20 3-20 3-20 3-21 3-21 3-21

Rev. 19

3-1


FAA APPROVED

TABLE OF CONTENTS(CONT)

Subject

Paragraph

COMMUNICATION RADIO FAILURE .................................... CABIN HEATER...................................3-11................. LANDING GEAR ........................................ STATIC PORT OBSTRUCTION ................................................. STICK CENTERING INDICATOR FAILURE ..............................

3-10-B ............

3-12................ 3-13................ 3-14................

Page Number 3-22 3-22 3-22 3-22 3-22

LIST OF TABLES

Table Number

Title

WARNING (RED) LIGHTS.......................................................... 3-1 .................. CAUTION (AMBER) LIGHTS...................................... 3-2.................. DIGITAL AFCS ERROR CODES ........................................... 3-3..................

3-2

Rev. 19

Page Number 3-23 3-25 3-32

FAA APPROVED

T N I * Section3 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 3-1. INTRODUCTION

The following procedures contain the indications of equipment or system failure or malfunction, the use of emergency features of primary and backup systems, and appropriate warnings, cautions, and explanatory notes. Table 3-1 lists fault conditions and corrective actions required for illumination of red warning lights. Table 3-2 addresses malfunction procedures associated with amber caution lights. Table 3-3 lists the Digital AFCS error codes.

BHT-412-FM-4

LAND AS SOON AS The duration of the PRACTICAL flight and landing site are at the discretion of the pilot. Extended flight beyond the nearest approved landing area is not recommended.

The following terms are used to describe the operating condition of a system, subsystem, assembly, or component.

All corrective action procedures listed herein assume the pilot gives first priority to aircraft control and a safe flight path.

Affected

Fails to operate in the intended or usual manner.

The helicopter should not be operated following any emergency landing or shutdown until the cause of the malfunction has been determined and corrective maintenance action taken.

Normal

Operates in the intended or usual manner.

3-3. ENGINE

3-2. DEFINITIONS

The following terms indicate the degree of urgency in landing the helicopter.

LANDAS SOON AS Land without delay at POSSIBLE the nearest suitable area (i.e., open field) at which a safe approach and landing is reasonably assured.

3-3-A. SINGLEENGINEFAILURE

The Ni rpm of the normally operating engine is allowed to droop to 97% during transition from twin engine operation to single engine operation. The N,, rpm should be increased to 100%if possible.

Flight can be continued on remaining engine until a desirable landing site is available. There are certain combinations of gross weight, altitude, and cold ambient temperatures at which a single engine approach will result in the OEI torque limit being exceeded. A run-on landing at 20 to 30 knots is recommended.

3-3

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O RPM COULD EXCESSIVELY.

CAUTION

RUN ON LANDINGSMAY RESULT IN ROLL OSCILLATIONS WHILE ON THE GROUND. IF THIS OCCURS, LOWERING COLLECTIVE FULLY DOWN OR DISENGAGING AP 1 AND AP 2 WILL STOP THE OSCILLATIONS.

Loss of an engine while hovering at high gross weight and extremely cold conditions will most likely result in exceeding the OEI torque limit. If an overtorque is observed or suspected, an appropriate log book entry shall be made. Refer to Performance charts in Section 4. NOTE

If an engine restart is to be refer to ENGINE attempted, RESTART IN FLIGHT in this Section. · INDICATION:

ENG 1 OUT or ENG 2 OUT warning light illuminated. N1rpm below 53% and decreasing.

DECAY

CAUTION

DURING COLD WEATHER OPERATIONS CAREFULLY MONITOR TORQUE OF THE NORMAL ENGINE WHEN ONE ENGINEFAILS OR IS SHUT DOWN IN FLIGHT.

Collective - Reduce as required to maintain rotor rpm and power within OEI limits. Airspeed - 70 KIAS or as required to maintain flight within OEI limits.

RPM switch - INCR; set N11 rpm at 100% if possible. Throttle (affected engine) -

Closed.

BOOST PUMP switch (affected engine)

-

OFF.

FUEL switch (affected engine) -

OFF.

FUEL XFEEDswitch -

OVRDCLOSE.

Fuel INTCONswitch -

OPEN.

N11rpm below 85% and decreasing. ITT below 400°C and decreasing.

GEN switch (affected engine) NON-ESNTLBUS -

No. 1 or 2 OIL PRESSURE, DC GENERATOR, and PART SEP OFF caution lights illuminated. Normal engine N11 rpm droops to 97%.

OFF.

As desired.

If No. 2 engine failed:

BATTERYBUS 2 switch -

OFF;

BATTERY BUS 1 switch -

ON.

PROCEDURE:

MASTERCAUTIONlight -

WARNING

IF CORRECTIVEACTION IS NOT INITIATED IMMEDIATELY,ROTOR 3-4

Rev.17

Reset.

Altitude - Descend below 5000 feet Hp (if practical) to prevent possible fuel starvation if remaining boost pump fails. Land as soon as practical.


BHT-412-FM-4

3-3-B. ENGINE RESTART IN FLIGHT

For battery start (nonassisted):

The conditions which would warrant an attempt to restart the engine would probably be a flameout, caused by a malfunction of the automatic mode of the fuel control unit. The decision to attempt an engine restart during flight is the pilots responsibility.

BATTERY BUS switch (affected engine) - ON.

BATTERY BUS switch (normal engine)OFF.

*PROCEDURE:

START switch starter limitations.

Engage;

Engine oil pressure -

observe

Indicating a rise.

CAUTION ...............

WHEN RESTARTING ENGINE IN MANUAL FUEL CONTROL MODE, CAREFULLY MONITOR ITT.

CAUTION

IF THE CAUSE OF ENGINE FAILURE IS OBVIOUSLY MECHANICAL, AS EVIDENCED BY ABNORMAL SOUNDS, DO NOT ATTEMPT A RESTART.

Throttle (affected engine) -

BOOST PUMP switch - ON.

GOV switch MANUAL.

START switch rpm.

Closed.

(affected engine)

FUEL switch (affected engine) FUEL XFEED switch -

Throttle - Open slowly at 12% N, rpm until ITT begins to rise. Do not open throttle further until ITT and N, rpm stabilize.

ON.

NORM.

(affected

engine)

GEN switch (affected engine) -

OFF.

-

-

Centered

at 55% N,

CAUTION

WHEN OPERATING IN MANUAL FUEL CONTROL MODE, MAKE SLOW, SMOOTH THROTTLE MOVEMENTS TO AVOID COMPRESSOR STALL, OVERTEMP, OVERSPEED, AND POSSIBLE DRIVE TRAIN DAMAGE. COORDINATE THROTTLE AND COLLECTIVE PITCH CHANGES TO AVOID OVERLOADING THE NORMAL ENGINE.

CAUTION

OEI PERFORMANCE CAN BE AFFECTED DURING GENERATOR ASSISTED START (WITH BOTH BATTERY SWITCHES ON).

Rev. 17

3-5

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Throttle - Increase slowly; adjust as required to control torque, ITT, and N, rpm.

No. 1 and 2 OIL PRESSURE, DC GENERATOR, and PART SEP OFF caution lights illuminated. PROCEDURE:

NOTE

If torque of affected engine is controlled slightly (approximately 4%) below torque of normal engine, rotor rpm will be governed within limits automatically by normal engine.

GEN switches (both) - ON.

BATTERYBUS 2 switch - ON.

WARNING

IF CORRECTIVEACTION IS NOT INITIATED IMMEDIATELY,ROTOR RPM COULD DECAY EXCESSIVELY.

Collective pitch - Reduce. Establish autorotative glide at 65 to 90 KIAS. NOTE

FUEL TRANS switch (affected engine)--

ON.

FUEL INTCONswitch -

NORM.

Land as soon as practical.

secure If restart was unsuccessful, affected engine as prescribed in Single Engine Failure procedure

--

3-3-C. DUAL ENGINEFAILURE * INDICATION:

ENG 1 OUT and ENG 2 OUT warning lights illuminated. RPM warning light illuminated.

AIRSPEED for best angle of glide in autorotation is 90 KIAS, and AIRSPEED for minimum rate of descent is 65 KIAS. Autorotational rate of descent is a function of AIRSPEEDand ROTOR RPM and is virtually unaffected by gross weight and density altitude.

Accomplish

autorotative

landing.--

If time permits before landing and a restart will not be attempted, proceed as follows: Throttles (both) - Closed.

BOOST PUMPswitches (both) - OFF. FUEL switches (both) - OFF.

FUEL TRANS switches (both) - OFF.

Rotor rpm audio on.

After landing, complete shutdown.

N, rpm below 53% and decreasing (both engines). Ni, rpm below 85% and decreasing (both engines). ITT below 400°C and decreasing (both engines).

3-6

Rev.9

I

FAA APPROVED


3-3-D. ENGINE UNDERSPEED NOTE

Normal deviations of ROTORRPM from the governed setting may occur when large collective changes are made but should not be confused with fuel control failure, unless a large steady-state TORQUEsplit occurs.

If there is a low power demand (less than single engine power available) at the time of low side failure, ROTOR RPM and ENG RPM (N,,) of the affected engine will decrease and stabilize at or slightly below the governed value. Torque, ITT, and GAS PROD RPM (N1)of the affected engine will also decrease. As ROTORRPM decreases, the normal engine will increase TORQUE output to assume the load. If power demand is near zero, there may not be a significant torque split.

If there is a high power demand (greater than single engine power available) at the time of low side failure, ROTOR RPM will decrease along with ENG RPM (N11), TORQUE,ITT, and GAS PRODRPM (N1)of the affected engine. As ROTOR RPM decreases, the normal engine will increase to maximum power to assume the load, causing significant increases in TORQUE, ITT, and GAS PROD RPM (N1),while ENG RPM (N11)will remain below the governed value. · INDICATION:

Low ENG RPM (N11and ROTOR RPM (possibly with RPM warning light and audio if power demand is in excess of single engine power available). TORQUE split (proportional to power demand).

Low GAS PROD RPM (N1), ITT, and TORQUE on affected engine. PROCEDURE:

WARNING

IF CORRECTIVEACTION IS NOT INITIATED IMMEDIATELY,ROTOR RPM COULD DECAY EXCESSIVELY.

Collective - Adjust as necessary to maintain ROTORRPM. Airspeed - 70 KIAS or as required to maintain flight within OEI limits. Affected engine - Identify.

Throttle frictions - Tighten on normal engine; reduce on affected engine. Throttle (affected engine) flight idle. GOV switch MANUAL.

(affected

Reduce to

engine)

-

CAUTION

WHEN OPERATING IN MANUAL FUEL CONTROL MODE, MAKE SLOW, SMOOTH THROTTLE MOVEMENTS TO AVOID COMPRESSOR STALL, OVERTEMP, OVERSPEED, AND POSSIBLE DRIVE TRAIN DAMAGE. COORDINATE THROTTLE AND COLLECTIVE PITCH CHANGES TO AVOID OVERLOADING THE NORMAL ENGINE.

Throttle (affected engine) - Increase slowly. Adjust throttle and collective as required to maintain TORQUEof affected engine slightly below TORQUEof normal engine. MASTERCAUTION light -

Reset.

Land as soon as practical. Rev.17

3-7

FAA APPROVED


3-3-E. ENGINEOVERSPEEDFUELCONTROL/GOVERNOR FAILURE NOTE

Normal deviations of ROTORRPM from the governed setting may occur when large collective changes are made but should not be confused with fuel control failure, unless a large steady-state TORQUEsplit occurs.

If there is a low power demand (less than single engine power available) at the time of high side failure, ROTOR RPM and ENG RPM (N,,) of the affected engine will increase considerably above the governed value. TORQUE,ITT, and GAS PRODUCER RPM (N,) of the affected engine will also increase. As ENG RPM (Nl,) and ROTOR RPM increase above the governed value, the normal engine will reduce power to keep itself from overspeeding, and will indicate significantly lower TORQUE, ITT, and GAS PROD RPM (N,) than the affected engine.

If there is a high power demand (greater than single engine power available) at the time of high side failure, ROTORRPM and ENG RPM (Ni) of the affected engine will surge initially, along with TORQUE, ITT, and GAS PROD RPM (N,). As ENG RPM (Ni,) and ROTOR RPM increase, the normal engine will reduce power to keep itself from overspeeding. The affected engine then tries to assume all the load, which is beyond its capability due to the high power demand. ENG RPM (N,,) of the affected engine (and ROTOR RPM) will then decrease and rejoin the ENG RPM (N,,) of the normal engine, stabilizing at or slightly above the governed value as the normal engine adjusts power output to share the load. INDICATION:

High ENG RPM (N,,) and ROTOR RPM, possibly with RPM warning light. 3-8

Rev.17

Definite TORQUE split (proportional to power demand).

High GAS PROD RPM (Ni), ITT, and TORQUE on affected engine.

Return of ENG RPM (Ni,) and ROTOR RPM to governed value (if power demand is high). *PROCEDURE:

CAUTION

IF CORRECTIVE ACTION IS NOT IMMEDIATELY, INITIATED ROTORRPM

CAN OVERSPEED

EXCESSIVELY.

Collective - Adjust as necessary to maintain ROTORRPM. Affected engine -

Identify.

Throttle (affected engine) - Reduce to maintain TORQUE at or slightly below TORQUEof normal engine. Throttle frictions - Tighten on normal engine; reduce on affected engine. Throttle (affected engine) idle. GOV switch MANUAL.

(affected

Reduce to

engine)

-

NCAUTION

WHEN OPERATING IN MANUAL FUEL CONTROL MODE, MAKE SLOW,SMOOTH, COORDINATED THROTTLE AND COLLECTIVE MOVEMENTS TO AVOID COMPRESSOR STALL, OVERTEMP, UNDERSPEED/ OVERSPEED, AND POSSIBLE DRIVETRAINDAMAGE.

I


FAA APPROVED

Throttle (affected engine) - increase slowly. Adjust throttle and collective as required to maintain TORQUEof affected engine slightly below TORQUEof normal engine. MASTERCAUTION- Reset.

IF NR OR MAST TORQUE EXCEEDS 110% OR OVERSPEED VALUES ARE UNKNOWN, LAND AS SOON AS POSSIBLE. Land as soon as practical.

3-3-E-1.

GOVERNOR ACTUATOR FAILURE (FULL INCREASE)

* INDICATION: N,i and rotor rpm increase to approximately 101%. RPM INCR-DECRswitch Inoperative.

PROCEDURE: If this failure occurs during takeoff or landing, no Immediate corrective action is necessary to complete either maneuver. As soon as practical, roll back both throttles to maintain 97 to 100%N,, rpm. Further adjustment of collective and throttles simultaneously will allow full power at pilots discretion. Land as soon as practical.

3-3-E-2.

GOVERNOR ACTUATOR FAILURE (FULL DECREASE) Published Flight Manual performance may not be attainable.

3-3-F. ENGINE OVERSPEEDDRIVESHAFT FAILURE

TRANSMISSIONWILL RESULT IN COMPLETELOSS OF POWERTO THE MAIN ROTOR. ALTHOUGH THE COCKPIT INDICATIONSFOR A DRIVESHAFT FAILURE ARE SOMEWHATCOMPARABLETO A DUAL ENGINE FAILURE, IT IS IMPERATIVE THAT AUTOROTATIVE FLIGHT PROCEDURESBE ESTABLISHED IMMEDIATELY. FAILURE TO REACT IMMEDIATELY TO THE LOW ROTORRPMAUDIO SIGNAL, CAUTION LIGHT, AND TACHOMETERINDICATION WILL RESULTIN LOSSOF CONTROL.

INDICATION:

Left yaw.

Rapid decrease in N Rapid increase in N,,.

Illumination of rotor RPM warning light with audio. Possible Increase in noise due to: Overspeeding engine turbines.

Overspeeding combining gearbox. Driveshaft breakage.

PROCEDURE:

Collective - As required to establish autorotative descent.

Airspeed - Establish airspeed for minimum rate of descent or maximum glide. Throttles -

Close, If time permits.

Controls landing.

As required for autorotative

3-3-G. ENGINE COMPRESSOR STALL * INDICATION: Engine 'pops'.

WARNING FAILURE OF THE MAIN DRIVESHAFT TO THE

High or erratic ITT. Decreasing or erratic N1 and N2 RPM. Torque oscillations. Rev.18

3-9


FAA APPROVED

*PROCEDURE:

Collective -

Reduce, maintain slow

flight

Heater switch -

INDICATION:

FIRE 1 PULL or FIRE 2 PULL handle

Reposition

ITT and N, -

Check for normal

indications

SEAT BELT and no smoking signs ON.

If an audible compressor

stall is noted,

refer to maintenance to determine cause.

*PROCEDURE:

Abort start of affected engine as follows: Throttle -

Closed

FUEL XFEED switch -

BOOST PUMP switch -

3-3-H. ENGINE HOT START/ SHUTDOWN

Appropriate FIREhandle -

A hot start is caused by a combination

of

excessive fuel in the combustion chamber and delayed fuel ignition. Internal and external damage can result.

Flames emitting from the exhaust. Excessive ITT indication.

FIRE EXT swtch -

Pull

MAIN.

FIRE warning light remains on more

* PROCEDURE:

FIRE EXT switch -

*PROCEDURE:

RESERVE

Complete engine shutdown

Abort start of affected engine as follows:

Throttle - Closed; keep starter engaged

BOOST PUMP switch -

Starter -

OFF

FUEL switch - OFF

* INDICATION:

FUEL switch -

OVRD CLOSE

OFF

OFF

Continue to energize until

ITT decreases.

Complete shutdown.

Exit helicopter and check for damage.

If ITT limits for starting were exceeded, refer to Engine Maintenance Manual for inspection requirements.

Exit helicopter.

3-4-A-2.

DURING TAKEOFF OR APPROACH APPROACH If a landing site is available, begin approach; if not available, continue takeoff.

* INDICATION

FIRE 1 PULL or FIRE 2 PULL handle

illuminated.

*PROCEDURE: Airspeed -

45 KIAS minimum.

3-4. FIRE

Collective

3-4-A. ENGINE FIRE

Appropriate FIRE PULL handle -

3-4-A-1.

FIRE EXT switch - MAIN.

3-10

DURING START Rev. 17

permitting)

-

Reduce (altitude

Pull


FAA APPROVED

If FIRE warning light remains on more than 10 seconds: PROCEDURE:

FIRE EXT switch N11rpm -

RESERVE

If reserve bottle has not been discharged, FIRE PULL handle (affected engine) - In (to provide fire protection for normal engine). GEN switch (affected engine) -

Set at 100%

Land as soon as possible

Exit helicopter.

IN FLIGHT


OFF


ON

NON ESNTL BUS switch as desired.

PROCEDURE:

MANUAL,

After landing, proceed as follows:

Initiate emergency descent immediately if possible.


Shut down affected engine as follows:

Exit helicopter.

Appropriate FIRE PULL handle -

Pull

Throttle (affected engine) - Closed FIRE EXT switch -

OFF.

If No. 2 engine was shut down:


3-4-A-3.

NOTE

MAIN

3-4-B. CABIN SMOKE OR FUMES INDICATION:

Smoke, toxic fumes, etc., in cabin.

FUEL XFEEDswitch - OVRD CLOSE

*PROCEDURE:

BOOST PUMPswitch (affected engine)

-

OFF


FUEL switch (affected engine) Fuel INTCONswitch -

OFF

OPEN.

If FIRE warning light remains on: FIRE EXT switch -

Vents and windows -

ON.

Open.

If additional ventilation is required: Airspeed less.

Reduce to 60 KIAS or

RESERVE

Passenger doors -

N11rpm (normal engine) -

Set at 100%

Land as soon as possible

If a landing site is not readily available, proceed as follows:

Open.

If time and altitude permit and the source is suspected to be electrical, attempt to identify and isolate the affected system. Land as soon as possible. Rev.17

3-11


BHT-412-FM-4

3-4-C. FIRE

BAGGAGE COMPARTMENT

3-5-A-1.

PROCEDURE:

INDICATION:

BAGGAGE illuminates.

LOSS OF T/R THRUST AT HOVER

FIRE warning

light

Close throttles immediately and make a hovering autorotation landing. A slight rotation can be expected on touchdown.

*PROCEDURE:

Reduce power to minimum required.

3-5-A-2.

CLIMB

Land as soon as possible.

The degree of right yaw upon failure will

Inspect tailboom area for damage.

be greater than that experienced

torque

The key to successful handling of a tail

rotor emergency lies in pilot ability to quickly recognize the type of malfunction and to select the proper emergency procedure. Following is a discussion of some types of tail rotor malfunctions

in level

flight due to the higher power and anti-

3-5. TAIL ROTOR

their probable effects.

LOSS OF T/R THRUST IN

and

3-5-A. COMPLETE LOSS OF TAIL ROTOR THRUST

settings.

*PROCEDURE:

Close throttles and lower collective pitch immediately. Establish a glide speed slightly above normal autorotation approach speed.

If a turn is required to reach a more desirable place to land or to align into the wind, make it to the right. A turn to the right can be more nearly streamlined by the use of a little power.

*INDICATION:

This is a situation involving a break in the

drive system,

such as a severed

Once aligned

controlled

for landing,

yaw can be

in the following

manner:

driveshaft, wherein the tail rotor stops

turning

and delivers no thrust. A failure of

RIGHT YAW

this type in powered flight will result in the nose of the helicopter swinging to the right (left side slip) and usually a roll of the fuselage. Nose down tucking will also be present. The severity of the initial

If the nose yaws right with power off, a pulse of up collective will produce more friction in the mast thrust bearings, creating a left moment. The greater the

density altitude, gross weight center of

response will be.

reaction will be affected by airspeed, gravity, and power being used.

input of the pulse, the more the

WARNING

DO NOT ALLOW ROTOR RPM TO DECAY BELOW MINIMUM LIMITS.

Moving the collective upward abruptly increases rotor loading. Do not hold 3-12

Rev. 17


FAA APPROVED

collective up, as rotor rpm will decrease lower than desirable. It is essential that the collective be returned to the down position for autorotation. This cycle is one pulse. The pulse should be rapid (up and down) but should not be used at low a I t i t u d e s. LEFT YAW

If the nose yaws left with power off, a slight addition of power should arrest it. Further increase in power results in more right yaw response. LANDING

CAUTION

RUN ON LANDINGS MAY RESULT IN ROLL OSCILLATIONS WHILE ON THE GROUND. IF THIS OCCURS, LOWERING COLLECTIVE FULLY DOWN OR DISENGAGING AP 1 AND AP 2 WILL STOP THE OSCILLATIONS.

During the final stages of the approach, a mild flare should be executed and all power to the rotor should be off. Maintain helicopter in a slight flare and use collective smoothly to execute a soft, slightly nose-high landing. Landing on the aft portion of the skids will tend to correct side drift. This technique will, in most cases, result In a run-on type landing. ZERO GROUNDSPEED LANDING

as near level as possible. A more positive and abrupt use of collective will also be necessary. CAUTION

THE FLARE AND THE ABRUPT USE OF COLLECTIVE WILL CAUSE THE NOSE TO ROTATE LEFT, BUT DO NOT CORRECT WITH THROTTLE. ALTHOUGH APPLICATION OF THROTTLE WILL RESULT IN ROTATION TO THE RIGHT,ADDITIONOF POWER IS A VERY STRONG RESPONSE MEASURE AND IS TOO SENSITIVE FOR THE PILOT TO MANAGE PROPERLY AT THIS TIME. DO NOT ADD POWER AT THIS TIME. SLIGHT ROTATION UPON TOUCHDOWN AT ZERO GROUND SPEED SHOULD NOT CAUSEANY REAL PROBLEM.

3-5-A-3.

LOSS OF T/R THRUST IN LEVEL FLIGHT OR POWER DIVE

*PROCEDURE:

Close throttles and reduce collective pitch immediately. Attain an airspeed slightly above the normal autorotative glide speed.

If altitude permits with airspeed above 60 knots, throttle and pitch may be gently applied to see if some degree of powered flight can be resumed. If any adverse yawing is experienced, re-enter autorotation and continue descent to a landing. The landing technique is the same as prescribed for the climb condition above.

If it is essential that the landing be made at zero ground speed, the only change in the technique described previously is that the flare should be executed closer to the ground and more steeply. This will require a more rapid forward cyclic input to land

Rev.17

3-13

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 3-5-A-4.

LOSS OF T/R THRUST DURING DESCENT (LOW POWER OR POWER OFF)

PROCEDURE:

If both throttles are not closed at the time of the failure, close them. Proceed as prescribed in the previous condition of level flight.

3-5-B-2.

FIXED PITCH FAILURE IN FLIGHT

If tail rotor fixed pitch failure occurs during climb (left pedal applied), cruise neutral pedals), and (approximately descent (right pedal applied), a descent and landing can be effected safely by use of power and throttle changes. *PROCEDURE:

3-5-A-5.

LOSS OF TAIL ROTOR COMPONENTS

The loss of any tail rotor components will result in a forward center of gravity shift. Other than additional nose down pitching, this situation would be quite similar to complete loss of tail rotor thrust, as discussed above.

3-5-B. TAIL ROTOR FIXED PITCH FAILURES INDICATION:

Tail rotor pitch change control failures are characterized either by a lack of directional response when a pedal is pushed or by locked pedals. If pedals cannot be moved with a moderate amount of force, do not attempt to apply a maximum effort, since a more serious malfunction could result.

3-5-B-1.

FIXED PITCH FAILURE AT HOVER

*PROCEDURE:

Do not close throttles unless a severe right yaw occurs. If pedals lock in any position at a hover, landing from a hover can be accomplished with greater safety under power controlled flight rather than and entering by closing throttles autorotation.

3-14

Rev.17

is in a trimmed If the helicopter condition when the malfunction is discovered, engine power and airspeed should be noted and the aircraft should be flown to a suitable landing area. Combinations of engine torque, rotor rpm, and airspeed will correct or aggravate yaw attitude and these should be adjusted as required to control yaw during landing.

RIGHT PEDAL LOCKED FORWARD OF NEUTRAL Power should be reduced and engine rpm maintained within the green arc. This will help streamline the helicopter in flight. Right turns are easier than left turns. Airspeed should be maintained at or above 60 knots.

Execute a shallow to normal approach, maintaining engine rpm and an airspeed of about 60 knots during the initial part of the approach. At 60 to 75 feet AGL and when the landing area can be made, start a slow deceleration to arrive at the intended landing point with about 25 knots indicated airspeed.

At 2 to 5 feet AGL, slowly reduce throttle to overcome yaw effect and allow helicopter to settle. When aligned with the landing area, allow aircraft to touch down.

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O CAUTION

RUN ON LANDINGSMAY RESULT IN ROLL OSCILLATIONS WHILE ON THE GROUND. IF THIS OCCURS, LOWERING COLLECTIVE FULLY DOWN OR DISENGAGING AP 1 AND AP 2 WILL STOP THE OSCILLATIONS.

After ground contact use collective and throttle as necessary to maintain alignment with landing strip, and to minimize forward speed. If helicopter starts to turn, move cyclic as necessary to follow the turn until helicopter comes to a complete stop.

LEFT PEDAL LOCKED FORWARD OF NEUTRAL

Reduce power and maintain engine rpm within the green arc. Normal turns can be safely made under these conditions, although the nose may be displaced to the left (depending on how far the pedal is forward). Maintain about 60 knots during the initial part of the approach. On final approach, reduce throttles to minimum operating rpm (97% NR) and simultaneously begin a slow deceleration so as to arrive at a point about two feet above the intended touchdown area as effective translational lift is lost.

Apply collective pitch (maintaining minimum operating rpm) to stop the rate of descent and forward speed, and to align the helicopter with the intended landing path. If not aligned after pitch application, increase throttle to further help with alignment. Allow helicopter to touch down at near-zero ground speed, maintaining alignment with throttle.

BHT-412-FM-4

Reduce power and maintain engine rpm within the green arc. Normal turns can be safely made under these conditions.

Execute a shallow approach, holding 60 knots airspeed during the initial part of the approach. At 50 to 75 feet AGL and when the landing area can be made, start a slow deceleration to arrive at the intended landing point with 25 knots indicated airspeed.

At 2 to 5 feet AGL, use throttle slowly as necessary to maintain alignment with the landing area and to control yaw; do not allow helicopter to settle until alignment is assured, then touch down. CAUTION

RUN ON LANDINGSMAY RESULT IN ROLL OSCILLATIONS WHILE ON THE GROUND. IF THIS OCCURS, LOWERING COLLECTIVE FULLY DOWN OR DISENGAGING AP 1 AND AP 2 WILL STOPTHE OSCILLATIONS.

After ground contact, use collective and throttle as necessary to minimize forward speed and to maintain alignment. Move cyclic as necessary to follow the turn until helicopter has come to a complete stop.

3-5-C. LOSS OF PITCHCHANGE CONTROLLINKAGE INDICATION:

In this type of failure, the pitch-change mechanism is broken at some point and

PEDALS LOCKEDIN NEUTRAL Rev.17

3-15

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O the tail rotor will assume a blade angle determined by the aerodynamic and counterbalance forces. PROCEDURE:

The corrective action procedures are described in FIXED PITCH FAILURES above. The specific procedure to be used depends on the yaw change experienced.

3-6. HYDRAULICSYSTEM

This helicopter has two independent hydraulic boost systems, both of which supply power to the flight control system for the main rotor. The tail rotor control system is powered by system 1 only.

If a hydraulic system failure occurs shortly after the helicopter has been cold soaked at or below -25°C (-13°F) and the center of gravity of the helicopter is near the forward or aft limit, some resistance may occur when the cyclic is near control position extremes. This resistance can be overcome by increased pilot effort.

WARNING

DO NOT EXTEND FLIGHT WITH FAILED HYDRAULIC SYSTEM. IS NOT THE HELICOPTER CONTROLLABLE WITH BOTH SYSTEMS HYDRAULIC INOPERATIVE.

COLD WEATHER DURING OPERATIONAVOID HIGH RATES OF CLIMB. MAKE APPROACHES AND LANDINGS INTO THE WIND. AVOID EXTENDED HOVERING AND DO NOT HOVER WITH THE WIND COMING FROM THE AFT LEFT QUADRANT.

Affected system -

Identify positively.

HYDR SYS switch (affected system)OFF. N,i rpm -

Set to 100%.

Reset.

MASTERCAUTIONlight -


INDICATION:

NO. 1 or NO. 2 HYDRAULICcaution light illuminates. Abnormal (low, high, or fluctuating) hydraulic pressure in affected system.

Possible high temperature in affected system.

Increased pedal forces (if system 1 failed). Increased cyclic forces near control extremes (cold weather only). *PROCEDURE:

If either hydraulic system fails, or if system temperature or pressure exceeds limits, proceed as follows: 3-16

Rev.17

3-7. ELECTRICALSYSTEM 3-7-A. DC POWERFAILURE INDICATION:

DC GENERATOR illuminates.

caution

light

All lighting and avionics on nonessential buses inoperative. *PROCEDURE:

GEN FIELD and GEN RESET circuit breakers -

Check in.

GEN switch (affected RESET,then ON.

generator)

-


BHT-412-FM-4

If generator remains proceed as follows:

inoperative,

GEN switch (affected generator) - OFF.

BATTERY BUS 1 switch - ON.

MANUAL.

Monitor; if load exceeds

NON-ESNTLBUS switch - As desired. Switch off unnecessary equipment as required.

If both generators fail and neither will reset, proceed as follows:

DO NOT SELECT EMERG LOAD

AT PRESSURE ALTITUDES ABOVE 5000 FEET. BOTH FUEL BOOST PUMPS WILL BECOME INOPERATIVE, RESULTING IN POSSIBLEFUEL STARVATION.

As desired.

A fully charged battery will provide sufficient electrical power for approximately 30 minutes with essential equipment operating normally except for limitations as follows:

Comm


3-7-B. AC POWER FAILURE * INDICATION:

NO. 1 or NO. 2 INVERTERcaution light illuminates.

Possible loss of power to certain ac instruments (with no INVERTERcaution light).

If either INVERTER caution illuminates, proceed as follows:

AC voltmeters - Check to determine that remaining inverter has assumed all ac loads. INV PWRcircuit breakers -

Check in.

If power Is lost only to certain ac instruments, but INVERTER caution lights remain out, proceed as follows:

AC FEEDERScircuit breakers (8 each)

-

#1 -

light

AP 1 or AP 2 button (affected system) - Depress to reengage autopilot.

NOTE

VHF

2 minutes.

*PROCEDURE:

CAUTION

Emergency load switch -

I

2 minutes.

With emergency load switch In EMERG LOAD position, battery will provide approximately 90 minutes of electrical power.

BATTERY BUS 2 switch - OFF.

DC AMPS limit:

Searchlight -

Windshield Wipers -

If No. 2 generator failed:

NON-ESNTLBUS switch -

Utility Light (overhead portable) 3 minutes.

Transmit

minutes, Receive 30 minutes. Pitot Static Heaters minutes.

Check in.

3

7 1/2

During IFR flight, if both inverters fail, land as soon as practical; or continue flight under VFR, if desired. Rev.18

3-17


BHT-412-FM-4

FUEL SYSTEM

3-8.

3-8-A.

FUEL BOOST PUMP FAILURE.

* INDICATION:

NO. 1 or NO. 2 FUEL BOOST caution light illuminates. PROCEDURE:

CAUTION

IF EITHER BOOST PUMP FAILS, BE WILL FUEL USABLE APPROXIMATELY 60 POUNDS LESS THAN INDICATED.

If practical, descend below 5000 feet Hp to prevent possible fuel starvation in the event remaining boost pump fails. FUEL INTCON switch -

If FUEL XFEED switch is in NORM position, crossfeed valve will be automatically opened allowing the remaining pump to supply fuel to both engines. Land as soon as practical.

FUEL FILTER PARTIALLY BLOCKED.

Land as soon as practical. 3-8-C.

Power loss to both signal conditioners will give a zero quantity indication. NOTE

A power failure to the signal conditioner will disable the FUEL LOW caution panel indication and alter the FUEL TRANS caution indication for affected fuel system. Refer to Table 3-2. * PROCEDURE:

FUEL QTY circuit breaker - Recycle. (Affected side.) FUEL INTCON switch - OPEN. NOTE

OPEN.

NOTE

3-8-B.

(Possible power failure to the respective fuel signal conditioner. The digital indicator will read half total quantity.)

FUEL QUANTITY INDICATIONS

MALFUNCTION

Allow sufficient time for fuel levels to equalize. Approximate fuel loads may be obtained by doubling remaining fuel quantity indicated.

3-9.

AUTOMATIC FLIGHT CONTROLS SYSTEM

Automatic Flight Control System (AFCS) can be affected by malfunctions of any vertical gyros, yaw rate gyros, (AHRS for S/N 36248 and subsequent), inverters, accelerometers (4 axis only), or by other electrical malfunctions. Failure of No. 1 hydraulic system will render AFCS yaw axis inoperative but will not affect pitch or roll SAS or ATT mode functions. Failure of No. 2 hydraulic system will not affect AFCS.

* INDICATION:

One fuel quantity indicator goes to zero from a previously normal condition.

3-18

Rev. 19

Digital AFCS fault codes will be displayed on the AL-300 Air Data Display. Table 3-3 lists error codes and descriptions.


BHT-412-FM-4

NOTE

If both autopilots are disengaged, the following procedures do not apply.

3-9-A.

AFCS FAILS TO ENGAGE OR

DISENGAGES

* INDICATION:

CAUTION

DO NOT ATTEMPT TO RESET ANY CIRCUIT BREAKER MORE THAN ONCE. INV 1 PWR and INV 2 PWR AC FEEDERS(8)

NO. 1 AUTOPILOT/NO. 2 AUTOPILOT caution light illuminated.

AP 1 or AP 2 off (light not illuminated).

Error code(s) displayed on AL-300. If more than one error exists, the AL-300 will continuously scroll through all error codes.

Possible erratic API indications on AP 1 or AP 2.

Possible ATT FAIL flag displayed on pilot or copilot ADI. Possible illumination of NO. 1 or NO. 2 INVERTER caution light. NOTE

If inverter 1 or 2 fails, AP 1 or AP 2 will disengage, but can be reengaged by depressing respective button on AFCS control panel. * PROCEDURE:

Airspeed - Reduce to 115 KIAS or less.

INV 1 and 2 switches - ON; check NO. 1 and NO. 2 INVERTER caution lights extinguished.

Pilot and copilot ADIs - Check ATT FAIL flags retracted, indicators functioning properly. Check the following circuit breakers in:

NO. 1 and NO. 2 ESNTL BUS FEEDERS (ON MAIN DC) AFCS (No. 1 and No. 2)

AFCS 26V (No. 1 and No. 2)


PILOT ATT SYS and CPLT ATT SYS

PILOT AHRS and COPILOT AHRS 28VDC (S/N 36248 and subsequent) No. 1/No. 2 ATTD - 26Vac Reset

Check/

AP 1 or AP 2 switch (affected system) Press to reengage.

If either autopilot will not reengage, or if abnormal control disturbance occurs, proceed as follows: Affected autopilot Airspeed KIAS.

Disengage.

Maintain at or below 115

If IFR, land as soon as practical; or continue flight under VFR, if desired.

If fault condition cannot be cleared, use error code displayed on AL-300 to determine failed axis. Refer to Table 3-3 for a list of error codes. If both autopilots fail to reengage, proceed as follows:


Rev. 19

3-19


BHT-412-FM-4

3-9-B. AFCS FAILS TO HOLD ATTITUDE

PROCEDURE:

. PROCEDURE:

FORCE TRIM switch - Check ON. SAS/ATT button illuminated.

Check ATT light

If AUTOTRIMcaution light illuminated Follow procedure for AFCS FAILS TO ENGAGEOR DISENGAGE(Paragraph 39-A).

3-9-C. AFCS HARDOVER OR ABNORMAL CONTROL DISTURBANCE WARNING

IF AP 1 OR AP 2 FAILS OR IS REDUCE DISENGAGED, AIRSPEEDTO 115 KIAS OR LESS. PROCEDURE:

Cyclic FORCE TRIM release switch Press.

Correct helicopter attitude with cyclic and pedals, then release cyclic FORCE TRIM SWITCH. Airspeed -

AL-300 will continuously scroll through all error codes.

Ensure cyclic friction set at minimum.

Attempt to reset AUTOTRIM light by cycling either AP 1 or AP 2 OFF, then ON.

If fault condition cannot be cleared, use error code displayed on AL-300 to determine failed axis. Refer to Table 3-3 for a list of error codes. NOTE

If AUTOTRIM caution light illuminates and no error codes are displayed on the AL-300 (ADD), then either or both autopilot computers may be in AOG mode. AOG mode inhibits the autotrim and attitude hold functions. This note does not apply to Flight Director equipped helicopters.

3-9-E. PITCH OR ROLL FAILURE

Pilot may choose to remain in ATT mode, using cyclic FORCE TRIM release switch and APIs to manually re-trim as required to keep actuators working near center; or pilot may select SAS mode with FORCE TRIM switch ON or OFF as desired.

Reduceto 115 KIAS or less.

Actuator position indicators - Check both systems. If any API shows or erratic maximum displacement operation of any actuator, switch affected autopilot OFF. If IFR, land as soon as practical, or continue flight under VFR, if desired.

3-9-D. AUTOTRIM FAILURE * INDICATION:

AUTOTRIMcaution light illuminates. Error code(s) displayed on AL-300 (ADD). If more than one error exists, the Rev.17

3-9-F. AUTOTRIM RUNAWAY

An autotrim runaway can occur only when both AP 1 and AP 2 are ON in ATT mode. INDICATION:

An autotrim runaway in flight will be evidenced by the cyclic stick being driven in a direction opposite to the actuator position Indications (AP 1 or AP 2). This condition occurs because the series actuators will be driven to limit authority to compensate for the autotrim runaway. When the actuators are saturated (on stops), the helicopter will respond to the runaway trim command; however, with both AP 1 and AP 2 operative, the autotrim


BHT-412-FM-4

will be cut off automatically two seconds after actuator saturation. · PROCEDURE:

Cyclic FORCE TRIM release button Depress to center actuators and retrim to desired attitude. Airspeed - Reduce to 115 KIAS or less.

- Check in/Reset.

Maintain unaccelerated straight and level flight for a minimum of 30 seconds for AHRS to realign. Verify AHRS alignment complete.

MANUAL AHRS HEADING SYNCHRONIZATION

AP 2 or AP 1 - OFF.

*PROCEDURE:

APIs - Monitor for proper operation.

Pilot/copilot AHRS - 28 Vdc circuit breaker - Check in.


3-9-G. AHRS - ALIGNMENT IN FLIGHT (S/N 36248 AND SUBSEQUENT) NOTE

During in-flight realignment in turbulent conditions, prolonged invalid heading data may result despite valid attitude data. Refer to Manual AHRS Heading Synchronization.

A power loss of 0.5 seconds or more will require realignment of AHRS.

Pilot/Copilot HSI -115 breaker - Check in.

Vac circuit

MAG/DG mode switch - While in straight and level flight, switch to DG position, then return to MAG position. Verify AHRS alignment complete.

AHRS FAILS TO REALIGN * INDICATIONS:

Failed systems do not realign.

HSI and ADI do not function properly. Either autopilot will not reengage.

Abnormal control disturbances.

* INDICATIONS:

ATT, RT, or HDG fail flag displayed on pilot or copilot ADI/HSI.

* PROCEDURE:

No. 1 or No. 2 AUTOPILOT- Disengaged - Caution light illuminated.

Affected autopilot -

Error Code(s) displayed on AL-300. If more than one error exists, the AL-300 will continuously scroll through all error codes.

If IFR, land as soon as practical; or continue under VFR.

Airspeed - at or below 115 KIAS.

* PROCEDURE: Airspeed -

reduce to 115 KIAS or less.


Verify Disengaged.

Unaffected autopilot -

Verify Engaged.

Pilot/Copilot AHRS 28 Vdc Circuit Breaker - Check in/Reset. No. 1/No. 2 ATTD 26 Vac Circuit Breaker

Verify engaged.

NOTE

Possible erratic API/HSI/ADI indication.

AUTOMATIC AHRS REALIGNMENT

Verify disengaged.

Unaffected autopilot -

If fault condition cannot be cleared, use error code display on AL-300 to determine failed axis. Refer to Table 3-3 for a list of error codes

3-10. COMMUNICATION SYSTEM 3-10-A.

INTERCOM FAILURE

* INDICATION: Weak or no reception in headsets. Rev. 19

3-21

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W 0 O E I F V NLY O FAA APPROVED

BHT-412-FM-4

OFF immediately.

* PROCEDURE:

HEATER switch -

Check headset connection.

CABIN HTR circuit breaker out, do not reset.

Verify volume and ICS controls set properly.

Check prior to next flight.

Cycle ICS circuit breaker out and in.

3-12.

For single pilot operations only with Emergency Communications Panel installed:

NOT APPLICABLE TO SKID GEAR HELICOPTER!

Plug headset into EMERGENCY COMM jack (above and behind pilot position).

3-13.

Select desired radio on copilot ICS panel. Key selected radio with EMERGENCY COMM switch (on center pedestal).

3-10-B.

COMMUNICATION RADIO FAILURE.

LANDING GEAR

STATIC PORT OBSTRUCTION

* INDICATION:

Erratic readings from the airspeed indicator, vertical speed indicator, and altimeter when operating helicopter in rain with STATIC SOURCE switch in PRI position.

* INDICATION:

* PROCEDURE:

Weak or no reception in headset.

Windows and vents - Close.

PROCEDURE:

Verify proper radio selected.

Verify volume properly adjusted.

Cycle appropriate circuit breaker out and back in.

3-11.

CABIN HEATER

A malfunction in the bleed air heater controls may or may not cause heater to become inoperative. * INDICATION:

HEATER AIR LINE caution light illuminates.

Heated airflow does not shut off when thermostat knob is turned to full cold position. * PROCEDURE:

3-22

Rev. 19

OFF.

HEATER switch -

STATIC SOURCE switch - ALTN.

Verify frequency properly set. Check headset connection.

Check; if

NOTE

This procedure selects an alternate static source (cabin air) for pilot side instruments only.

3-14.

STICK CENTERING INDICATOR FAILURE

* INDICATION:

CYC CTR caution lights fail to illuminate when cyclic is displaced 1.5 inches or more from center position while RPM warning light is illuminated. * PROCEDURE:

Maintain rotor rpm between 97 and 100% NR for ground operation until shutdown.


BHT-412-FM-4

Table 3-1.

PANEL WORDING

WARNING (RED) LIGHTS

FAULT CONDITION

CORRECTIVE ACTION

FIRE PULL (1 or 2)

Fire indication in No. 1 or Pull illuminated FIRE PULL handle. No. 2 engine compartment Select MAIN fire extinguisher. Close throttle of affected engine. Select RESERVE fire extinguisher if necessary. Land as soon as possible.

BAGGAGEFIRE

Smoke in compartment

ENG OUT (1 or 2)

GAS PROD abnormally low Check ENG TORQUE, GAS PROD (below 53 ± 2%) RPM (N1) RPM (N1), ENG RPM, (N11),and ITT. on No. 1 or No. 2 engine. Adjust power and airspeed (70 KIAS). Reset remaining engine N11 to normal range. Close throttle of affected engine. Refer to ENGINE FAILURESand ENGINERESTARTIN FLIGHT procedures. Land as soon as practical.

XMSN PRESSURE

C BOX TEMP

Reduce power to minimum required. Land as soon as possible. Inspect tailboom area for damage.

O I L Transmission oil pressure Reduce power. Land as soon as below limit, possible.

XMSNOIL TEMP

C BOX PRESSURE

baggage

Transmission oil temperature above limit.

Reduce power. Check XMSN OIL temperature. If not within limits, land as soon as possible.

O I L Combining gearbox oil pressure below normal.

Reduce power. Land as soon as possible.

Combining gearbox oil temperature above limit.

Reduce power. Check temperature. If not within limits, land as soon as possible.

Rev.17

3-23


BHT-412-FM-4

Table 3-1. WARNING (RED) LIGHTS (Cont)

PANEL WORDING

BATTERYTEMP

FAULT CONDITION

CORRECTIVE ACTION

Battery case temperature above limit.

BATTERY switch soon as practical.

OFF. Land as

WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATIONOF BATTERY TEMP LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'SINSTRUCTIONS PRIOR TO RETURNTO SERVICE.

ROTORBRAKE

Rotor brake linings retracted.

RPM (with audio) or RPM (without audio)

Rotor rpm at or below 95%. Adjust collective pitch and/or RPM INCR-DECR switch as required. or Refer to Engine Fuel Control Rotor rpm at or above Malfunction procedures. 105%.

3-24

not

Check rotor brake handle fully up in detent. If light remains on, land as soon as possible.


FAA APPROVED

Table 3-2.

CAUTION (AMBER) LIGHTS

PANEL WORDING

FAULT CONDITION

OIL PRESSURE (ENG 1 or 2)

Engine oil pressure limit.

DC GENERATOR (ENG 1 or 2)

Failure of dc generator.

CORRECTIVE

below

ACTION

Shut down affected engine. INTCON switch - OPEN.

FUEL


GEN FIELD and GEN RESET circuit breakers - Check in. GEN switch (affected generator) - RESET, then ON. If light remains on, turn GEN switch OFF. If No. 2 generator failed: BATTERY BUS 2 switch BATTERY BUS 1 switch If

bus nonessential required: BUS NON-ESNTL MANUAL. DC AMPS - Monitor.

OFF. ON.

power

is

switch

If both generators fail:

CAUTION

DO NOT SELECT EMERG LOAD AT PRESSURE ALTITUDES ABOVE 5000 FEET. BOTH FUEL BOOST PUMPS WILL BECOME INOPERATIVE, RESULTING IN POSSIBLE FUEL STARVATION. Emergency load switch - As required. Land as soon as practical.

PART SEP OFF (ENG 1 or2)

Particle separator bypass door closed or circuit breaker out. Ice and dust protection system inoperative.

Check ENG 1 (or 2) RPM and PART SEP circuit breakers in.

Move PART SEP switch ON.

to OVRD


BHT-412-FM-4

Table 3-2.

PANEL WORDING

CAUTION (AMBER) LIGHTS (Cont)

FAULT CONDITION

CORRECTIVE ACTION

NO. 1 FUEL BOOST/ Fuel boost pump failure has If practical, descend below 5000 feet fuel possible NO. 2 FUEL BOOST occurred. Hp to prevent starvation If other boost pump tails. FUEL INTCONswitch - OPEN.

CAUTION

IF EITHER BOOST PUMP FAILS, BE FUEL WILL USABLE APPROXIMATELY60 POUNDSLESS THAN INDICATED.

NOTE

If either fuel boost pump fails, and the FUEL XFEED switch Is In NORM position, the crossfeed valve is opened automatically by a pressure switch, allowing either boost pump to furnish fuel to both engines.


NO. 1 FUEL FILTER/ Fuel filter NO. 2 FUEL FILTER blocked.

3-26

is partially



BHT-412-FM-4

Table 3-2. CAUTION (AMBER) LIGHTS (Cont)

PANEL WORDING

FAULT CONDITION

CORRECTIVE ACTION

FUEL LOW Fuel level in left or right Land as soon as practical. (Less than 100 lbs. cells at or below 190 difference between pounds. No. 1 and No. 2 fuel NOTE quantities)

I

Interconnect valve will open automatically when fuel level in The FUEL LOW light will opposite side decreases to 190 not illuminate for the pounds (as indicated by illumination affected side when fuel of FUEL INTCONcaution light). This quantity indication will allow the fuel quantity in the malfunction occurs. Refer lower cells to equalize. In the event to FUEL QUANTITY of a fuel boost pump failure, all fuel INDICATION in the lower cells will be available to MALFUNCTION. both engines through either boost pump. NOTE

FUEL INTCON caution light can be extinguished by placing FUEL INTCONswitch to OPEN position.

FUEL LOW (100 lbs. or more difference between No. 1 and No. 2 fuel quantities)

Possible fuel leak in cells with lower quantity.

FUEL INTCON switch CLOSE.

OVRD


GOV MANUAL (ENG Engine governor in manual Torque, ITT, and rpm must be 1 or 2) mode. controlled with throttle.

ENGINECHIP (ENG 1 Metal particles in engine or 2) oil.

Reduce power and shut down engine as soon as practical to minimize engine damage. Land as soon as practical.

FUEL VALVE (ENG 1

Check FUEL VALVE circuit breakers In. Land as soon as practical. If on ground, cycle FUEL switch.

Fuel

valve

not

properly

or 2)

seated or circuit out.

NO. 1 GEN OVHT/ NO. 2 GEN OVHT (if installed)

Generator overheating.

GEN switch -

CAUTION PANEL

Caution panel inoperative.

Check MASTER CAUTION circuit breaker In. Monitor aircraft instruments. Land as soon as practical.

breaker

OFF.

Rev.9

3-27


FAA APPROVED


PANEL WORDING

FAULT CONDITION

NO. 1 INVERTER/NO. Failure of 2 INVERTER Inverter;

ac

CORRECTIVE ACTION

power Check both ac voltmeters to determine that remaining Inverter automatically assumed load for failed inverter.

Check INV PWR circuit breakers in. Reengage AP 1 or AP 2. During IFR flight, if both inverters fall, land as soon as practical; or continue flight under VFR, If desired.

or

NO. 2 INVERTER

Emergency load switch in Place emergency load switch in EMERGLOAD position. NORMAL position, if electrical load shedding is not required.

EXTERNALPOWER

External power receptacle Check external power disconnected. door open. Check doors closed.

DOORLOCK

Passenger door(s) or Check doors closed and locked. baggage compartment door not locked.

BATTERY

Both BATTERY switches/ Turn one BATTERYswitch ON, other relays in the same position. OFF. If light remains on, reverse BATTERYswitch positions.

C BOX CHIP

Metal particle in combining Reduce power. gearbox oil. Land as soon as practical.

XMSNCHIP

Metal particles in Reduce Power. transmission oil (one or Land as soon as practical. more remote XMSN CHIP Indicators tripped).

42/90 BOX CHIP

Metal particles in 420or 90° Land as soon as practical. gearbox oil.

NO. 1 HYDRAULIC/ Hydraulic pressure below Verify fault and affected system NO. 2 HYDRAULIC limit or temperature above from gage readings. Turn off limit. affected system. Land as soon as possible.

3-28


BHT-412-FM-4

Table 3-2.

PANEL WORDING

NO. 1 FUEL TRANS/ NO. 2 FUEL TRANS

CAUTION (AMBER) LIGHTS (Cont)

FAULT CONDITION

CORRECTIVE

Fuel transfer pump or ejector pump malfunction (no fuel transfer from lower forward and middle cells to lower aft cell); or Check valve malfunction allowing fuel to leak from aft to mid cell after normal transfer is complete (total fuel 800 pounds or less).

Check FUEL TRANS circuit breaker - in. Check FUEL TRANS switch ON.

ACTION

NOTE

FUEL TRANS light will remain Illuminated after fuel transfer with fuel quantity indication malfunction. Refer to FUEL QUANTITY INDICATION MALFUNCTION.

CAUTION

IF EITHER TRANSFER PUMP FAILS, USABLE FUEL WILL BE 25 POUNDS LESS THAN INDICATED. If light remains illuminated: FUEL TRANS switch - OFF.

CAUTION

FUEL TRAPPED IN MID CELL IS UNUSABLE AND MUST BE SUBTRACTED FROM TOTAL FUEL QUANTITY INDICATION. Monitor MID TANK quantity periodically. Land as soon as Practical.

3-29


BHT-412-FM-4


PANEL WORDING

FAULT CONDITION CORRECTIVE ACTION FUEL INTCON Fuel interconnect valve not Check FUEL INTCON circuit (Switch In NORM fully closed. (Automatic breakers (both) in. FUEL INTCON position). valve opening is normal if switch - OPEN,then NORM. FUEL LOW light is also illuminated.) FUEL INTCON FUEL Interconnect valve Check FUEL INTCON circuit (Switch in OPEN not fully open or FUEL breakers in. FUELINTCONswitch position). INTCON circuit breakers OVRD CLOSE,then OPEN. out. FUEL XFEED

Fuel crossfeed valve not Check FUEL XFEED circuit breakers fully open or closed, or (both) in. Cycle FUELXFEEDswitch. FUEL XFEED circuit breakers out.

HEATER AIR LINE

Heater mixing malfunction.

NO. 1 AUTOPILOT/ NO. 2 AUTOPILOT

Autopilot 1 or 2 inoperative.

Engage affected autopilot. If unable to engage, continue flight on remaining system.

Automatic flight control system hardover; or Loss of ac power to AP 1 or AP 2; or Loss of attitude gyro input to AP 1 or AP 2 (possible disengagement of either or both autopilots.) or Auto trim malfunction. Displacement between AP 1 and AP 2 actuators at least 50 percent travel.

Reduce airspeed to 115 KIAS or below. Check AFCS control panel. If either autopilot is off, attempt to switch ON. (Refer to AFCS malfunction procedures.) During IFR flight, if both AP 1 and AP 2 are failed and will not reset, land as soon as practical; or continue flight under VFR, if desired. Reduce airspeed to 115 KIAS or below. Check actuator position panel. If APIs are centered, depress SYS 2 button to check AP 2 actuator displacement. Turn off affected system.

AUTOTRIM

Pitch, roll, inoperative.

or yaw trim

Cycle either AP 1 or AP 2 OFF, then ON. If light does not extinguish, follow AUTOTRIM FAILURE procedures in this section.

FDR SYS FAIL (if installed)

Flight data inoperative.

recorder

Refer to applicable regulations for operating with flight data recorder Inoperative.

3-30

valve

Turn HEATER immediately.

switch

OFF


BHT-412-FM-4


PANEL

WORDING

FAULT CONDITION

CORRECTIVE ACTION

FT OFF

Force trim inoperative.

Check FORCE TRIM switch ON and FORCE TRIM circuit breaker in. During IFR flight, If system remains inoperative, land as soon as practical; or continue flight under VFR, if desired. Pilot may increase cyclic friction to provide additional cyclic stabilization.

CYC CTR

Cyclic not centered.

Center cyclic.

OVER TORQ

Mast torque exceeds 100.5%.

Reduce power or severity of maneuver. Land as soon as practical.

AHRS FAN

AHRS fan inoperative

Avoid prolonged ground run operations if OAT exceeds 40 °C.

Rev. 19

3-31


BHT-412-FM-4

Table 3-3.

DIGITAL AFCS ERROR CODES

PWR UP

PRE FLT

CONT TEST

AL-300 DISPLAY

ERROR CODE DESCRIPTION

X X X X X X

X X

X X

X X X X

X X X X

EO E1 E2 E3 E4 E5 E6 E7 E8 E9

LOST PROCESSOR RAM PARTITION FAIL INSTRUCTION SET FAIL REGISTER TEST FAIL RAM RD/WR FAIL EPROM CHECKSUM FAIL CPU FAIL D/A-A/D FAIL INTERRUPT REG FAIL +15V MONITOR FAIL (POSSIBLE HARDWARE E/A FAIL)

X X X

X

E10

X X X X X X X

Ell E12 E13 E14 E15 E16 E17 E18 E19 E20 E21 E25 E26 E28 E29 E31 E32 E34 E35 E36 E38 E39

+5V MONITOR FAIL +10V MONITOR FAIL LVC MONITOR FAIL SAWTOOTH MONITOR FAIL ACTUATOR REFERENCE MONITOR FAIL WATCHDOG MONITOR FAIL TICKET FAILURE LVC FAIL SAWTOOTH FAIL ACTUATOR REFERENCE FAIL WATCHDOG FAIL TRIM END AROUND FAIL TRMMON CKSUM FAIL NO ROLL TRIM NO PITCH TRIM ROLL TRIM FAIL PITCH TRIM FAIL ROLL TRIM RUNAWAY PITCH TRIM RUNAWAY TRMRUN CKSUM FAIL YAW SERIES SERVO FAIL ROLL SERIES SERVO FAIL

X

E42 E55

X

E56

X

X X X X X X X X X X X X X X X X

X

PITCHSERIESSERVOFAIL

X

E101

VG MISCOMPARE ERROR CODE STORAGE ALERT:< 100 AVAILABLE ERROR CODE STORAGE ALERT: NONE AVAILABLE PITCH ARL TEST FAIL

X X X X X

E103 E104 E105 E106 E110

PITCH ACT MOVED DURING ARL TEST ROLL ACT MOVED DURING ARL TEST PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED NO ROLL TRIM

X

X

X

X 3-32

E40

-15V MONITORFAIL

E102

E111 E113

ROLL ARL TEST FAIL

NO PITCHTRIM

ROLL TRIM BACKWARDS


FAA APPROVED

Table 3-3.

PWR UP

DIGITAL AFCS ERROR CODES (Cont) AL-300 DISPLAY


X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X

E114 E115 E116 E117 E118 E119 E121 E124 E125 E126 E129 E131 E133 E134 E136 E137 E139 E141 E142 E143 E144 E147 E148 E149 E150 E151 E152 E155 E156 E157 E201 E204 E208 E209 E210

X X X X X X X X X X X X

E211 E212 E213 E216 E217 E219 E221 E222 E225 E226 E229 E231

PITCH TRIM BACKWARDS SERVO VALID DELAY (2 SEC) FAIL WATCHDOG CANNOT DISENGAGE AP YAW ACT/SERVO AMP FAIL ROLL ACT/SERVO AMP FAIL PITCH ACT/SERVO AMP FAIL FD ANN FAIL FD ANN FAIL OTHER POD DURING TEST ROD DURING TEST FTR DURING TEST TRIM SWITCH OFF DURING TEST PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED YAW SERIES ACT MOVED PITCH TRIM (H) E/A ON PITCH TRIM (L) E/A ON ROLL TRIM (H) E/A ON ROLL TRIM (L) E/A ON TRIM ENABLE FAIL TRIM ONE-SHOT FAIL PITCH TRIM (H) E/A OFF PITCH TRIM (L) E/A OFF ROLL TRIM (H) E/A OFF ROLL TRIM (L) E/A OFF SERVO VALID FAIL TRIM END AROUND FAIL APON OTHER INPUT FAIL AL-300 KNOB MOTION ANALOG GROUND FAIL ROLL ATTITUDE MORE THAN 10 DEG PITCH ATTITUDE MORE THAN 10 DEG YAW RATE GYRO MORE THAN 10 DEG/ SEC DIRECTIONAL GYRO FAIL VERTICAL GYRO NOT VALID DIRECTIONAL GYRO NOT VALID SERIAL I/O NOT VALID OTHER AUTOPILOT NOT IN TEST AP CONTROLLER BUTTON(S) PRESSED HARDWARE NOT VALID FOUR-WAY BEEP SWITCH PRESSED POD DURING TEST ROD DURING TEST FTR DURING TEST TRIM SWITCH OFF DURING TEST

PRE FLT

CONT TEST

3-33


BHT-412-FM-4

Table 3-3.

PWR UP

3-34

DIGITAL AFCS ERROR CODES (Cont) AL-300 DISPLAY


X

E233

ENG TORQUELESS THAN 5% OR GREATERTHAN 30%

X

E234

ENG TEMPLESS THAN 150 DEG OR GREATERTHAN 900 DEG C

X

E236

AIRSPEEDINPUTOPEN

X

E237

HEADINGNOT VALID

X

E241

X

E244

X

E245

C/A RESETBUTTONPRESSED ERASE EEPROMSWITCHON DURING TEST READ EEPROMSWITCHON DURING TEST

PRE FLT

CONT TEST

FAA APPROVED

BHT-412-FM-4

T * Section 4 N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TABLE OF CONTENTS

Page Number

Paragraph

Subject

INTRODUCTION ......................................................................... 4-1.................. POWER ASSURANCE CHECK ...................................... 4-2 .................. DENSITY ALTITUDE................................................................... 4-3.................. 4-4.................. HEIGHT -VELOCITY ENVELOPE.............................................. 4-5.................. HOVER CEILING......................................................................... 4-5-A.............. HOVER CEILING - IGE ........................................................ 4-5-B.............. HOVER CEILING - OGE ...................................................... 4-6.................. TAKEOFF DISTANCE................................................................. CLIMB AND DESCENT............................................................... 4-7.................. TWIN ENGINE RATE OF CLIMB............................................ 4-7-A.............. 4-7-B.............. SINGLE ENGINE RATE OF CLIMB........................................ 4-8 .................. AIRSPEED CALIBRATION ................................ 4-9.................. LANDING DISTANCE ................................................................. NOISE LEVELS..................................................4-10........ 4-10 .. . FAA NOISE CERTIFICATION ................................ 4-10-A............ 44-10-B............ ICAO ANNEX 16 NOISE CERTIFICATION ............................

4-3 4-3 4-3 4-4 4-4 4-4 4-4 4-4 4-4 4-4 4-5 4-5 4-5 4-5 4-5 4-6

LIST OF FIGURES

Figure Number

Title

Power assurance check ............................................................ Density altitude ......................................................................... Critical relative wind azimuths ................................................. Hover ceiling IGE ....................................................................... Hover ceiling OGE ........................................ Takeoff distance......................................................................... Twin engine rate of climb.......................................................... Single engine rate of climb ....................................................... ......... Airspeed calibration ........................................ .. Single engine landing distance ......................................

Page Number

4-1.................. 4-2 .................. 4-3.................. 4-3A............. 4-4.................. 4-5 .................. 4-6 .................. 4-7.................. 4-8 .................. 4-9...................

SEP 052000

Rev.21

4-7 4-9 4-10 4-10A/4-10B 4-11 4-23 4-25 4-49 4-55 4-56

4-1/4-2

FAA APPROVED

T * Section 4 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FM-4

4-1. INTRODUCTION

4-3. DENSITYALTITUDE

The performance data presented herein are derived from the engine manufacturer's specification power for the engine less installation losses. These data are applicable to the basic helicopter without any optional equipment which would appreciably affect lift, drag, or power available.

A Density Altitude Chart (figure 4-2) is provided to aid in calculation of performance and limitations. Density altitude is an expression of the density of the air in terms of height above sea level; hence, the less dense the air, the higher the density altitude. For standard conditions of temperature and pressure, density altitude is the same as pressure altitude. As temperature increases above standard for any altitude, the density altitude will also increase to values higher than pressure altitude. The chart expresses density altitude as a function of pressure altitude and temperature.

4-2. POWERASSURANCE CHECK

Power Assurance Check charts (figure 4-1) are provided to determine if the engines can produce installed specification power.

A power assurance check should be performed daily. Additional checks should be made if unusual operating conditions or indications arise. The hover check is performed prior to takeoff, and the in-flight check is provided for periodic in-flight monitoring of engine performance. Either power assurance check method may be selected at the discretion of the pilot. It is the pilots responsibility to accomplish the procedure safely, considering passenger load, terrain being overflown, and the qualifications of persons on board to assist in watching for other air traffic and to record power check data. If either engine does not meet the requirements of the hover or the in-flight power assurance check, published performance may not be achievable. The cause of engine power loss, or excessive ITT should be determined as soon as practical. Refer to Engine Maintenance Manual.

The chart also includes the inverse of the square root of the density ratio (1/1o), which is used to calculate KTAS by the relation:

EXAMPLE

If the ambient temperature is -15°C and the pressure altitude is 6000 feet, find the density altitude, 1//o, and true airspeed for 100 KCAS. SOLUTION

Enter the bottom of the chart at -15°C.

Move vertically upward to the 6000 foot pressure altitude line.

From this point, move horizontally to the left and read a density altitude of 4000 feet and move horizontally to the right and read 1/io equals 1.06. Rev.2

4-3

FAA APPROVED


True airspeed = KCAS x 1/Ja = 100 x 1.06 = 106 KTAS.

4-4. HEIGHT- VELOCITY ENVELOPE Refer to Section 1 - Limitations.

4-5.

HOVER CEILING

cyclic and directional control have been demonstrated in relative winds of 35 knots from any direction at or below 3000 feet HD. Improved control margins will be achieved by avoiding winds in the critical relative wind azimuth areas (figure 4-3). AREA B (shaded area) as shown on hover ceiling charts presents additional hover performance which can be achieved in calm winds or winds outside the critical relative wind azimuth areas.

4-5-A. HOVERCEILING- IGE

Adequate cyclic and directional control are available at the gross weights allowed by the Hover Ceiling IGE charts in relative winds up to 35 knots from any direction at or below 3000 feet HD. Improved control margins will be achieved by avoiding winds in the critical relative wind azimuth areas (figure 4-3). The Hover Ceiling chart (figure 4-3A) provides the maximum allowable gross weights for hovering IGE at all pressure altitude and outside air temperature with heater on or off. conditions Conversely, the hover ceiling altitude can be determined for any given gross weight.

4-5-B. HOVERCEILING- OGE

The Hover Ceiling charts (figure 4-4) provide maximum weights for hovering OGE at all pressure altitude and outside air temperature conditions with heater on or off. CAUTION

OGE HOVER OPERATION MAY OF RESULT IN VIOLATION LIMITATIONS. HEIGHT-VELOCITY

Some of the OGE hover ceiling charts are divided into two areas as follows:

AREA A (unshaded area) as shown on the hover ceiling charts presents hover performance for which satisfactory 4-4

Rev.17

NOTE

Tail rotor or cyclic control margin may preclude operation in AREA B of the hover ceiling charts when wind is in the the relative respective critical wind azimuth area.

4-6. TAKEOFF DISTANCE

The Takeoff Distance charts (figure 4-5) provide takeoff distances required to clear a 50 foot or 15 meter obstacle in a zero wind condition, using a takeoff flight path which will avoid the critical areas of the diagram (Section 1). Height-Velocity Takeoff is initiated from a hover at 4 feet (1.2 meters) skid height with climbout speed of 45 knots. NOTE

are not Downwind takeoffs because the recommended distance takeoff published performance cannot be achieved.

4-7.

CLIMB AND DESCENT

4-7-A. TWIN ENGINERATEOF CLIMB

The Twin Engine Rate of Climb charts (figure 4-6) provide the rates of climb that I can be obtained at all outside air temperatures/pressure altitudes/gross

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O weight combinations with heater on or off at maximum continuous power and takeoff power. NOTE

All rate of climb data are based on changes in true altitude (pressure altitude corrected for nonstandard temperature).

4-7-B. SINGLEENGINERATEOF CLIMB

The Single engine rate of climb charts I (Figure 4-7) provide the rates of climb that can be obtained at all outside air altitudes/gross temperatures/pressure weight combinations with heater off at maximum continuous power and 30 minute OEI power. NOTE

Published single engine performance is intended for emergency use only when one engine becomes inoperative due to an actual malfunction. Routine operation in excess of published twin engine operating limits can affect engine service life.

4-8. AIRSPEED CALIBRATION

The Airspeed calibration chart (Figure 4-8) provides calibrated airspeeds for all indicated airspeeds during level flight, climb and autorotation.

4-9. LANDINGDISTANCE

The Single engine landing distance chart I (Figure 4-9) provides the landing distances required to clear a 50 foot (15 meter) obstacle for all outside air temperatures, pressure altitudes, and gross weights. Landing distances are based on an approach condition of 45 KIAS and 500

feet per minute rate of descent and zero wind.

4-10. NOISE LEVELS 4-10-A. FAA NOISE CERTIFICATION

This helicopter is certified as a Stage 2 helicopter as prescribed in FAR Part 36, Subpart H, for gross weights up to and including the certificated maximum takeoff and landing weight of 11,900 pounds (5398 kilograms). There are no operating limitations in meeting the takeoff, flyover, or approach noise requirements. The following noise levels comply with FAR Part 36, Appendix H, Stage 2 noise level requirements. They were obtained by analysis of approved data from noise tests conducted under the provisions of FAR Part 36, Amendment 36-14. The certified noise levels are: Flight Condition

EPNL(EPNdB)

Takeoff

92.8

Flyover

93.4

Approach

95.6

NOTE

No determination has been made by the Federal Aviation Administration that the noise levels of this helicopter are or should be acceptable or unacceptable for operation at, into, or out of any airport.

VHis defined as the airspeed in level flight obtained using the minimum specification engine torque corresponding to maximum continuous power available for sea level 25°C ambient conditions at the relevant maximum certificated weight. The value of VH thus defined for this helicopter is 122 KTAS. Rev. 17

4-5


BHT-412-FM-4

4-10-B. ICAO ANNEX 16 NOISE CERTIFICATION

This helicopter complies with noise emission standards applicable to the aircraft as set out by International Civil Aviation Organization (ICAO) in Annex 16, Volume 1, Chapter 8, for gross weights up to and including certificated maximum takeoff and landing weight of 11,900 pounds (5398) kilograms). There are no operating limitations in meeting takeoff, flyover or approach noise requirements.

Following noise levels comply with ICAO ANNEX 16, Volume 1, Chapter 8 noise level

4-6

Rev.17

requirements. They were obtained by analysis of approved data from noise tests conducted under provisions of ICAO Annex 16, Volume 1, Second Edition 1988. Noise levels for the 412EP are: Flight Condition Takeoff Flyover Approach

EPNL(EPNdB) 92.8 93.4 95.6


BHT-412-FM-4

MODEL412

POWER ASSURANCE CHECK- HOVER PTBT-3DENGINE

HEATER/ECU- OFF.

IF INDICATEDITT EXCEEDSMAX ALLOWABLE,REPEAT CHECK STABILIZINGPOWERFOUR MINUTES.

THROTTLES: TEST ENGINE-FULL OPEN,FRICTIONED. OTHERENGINE-FLIGHT IDLE.

REPEAT CHECK USINGOTHERENGINE.

NIl RPM-97%.

COLLECTIVEPITCH-INCREASE UNTIL LIGHT ON SKIDSOR HOVERING.DO NOT EXCEED810 ITT, 103.1%NI RPM, OR 73.2% TORQUE.

IF EITHERENGINEEXCEEDSALLOWABLE ITT AFTER STABILIZING FOUR MINUTES,PUBLISHEDPERFORMANCE MAY NOT BE ACHIEVABLE.CAUSE SHOULDBE DETERMINEDAS SOON AS PRACTICAL.

STABILIZE POWERONE MINUTE,THEN RECORDPRESSURE ALTITUDE,OAT, ENGINETORQUE AND ITT.

ENTER CHART AT INDICATEDENGINETORQUE,MOVE UP TO INTERSECTPRESSUREALTITUDE, PROCEEDTO THE RIGHT TO INTERSECTOUTSIDEAIR TEMPERATURE, THEN MOVE UP TO READ VALUE FOR MAXIMUMALLOWABLEITT.

MAXIMUMALLOWABLEITT-°C

40

50

500

550

60

70

600

650

700

750

800

ENGINETORQUE-PERCENT (INDICATED)

412EP-FM-4-1-I

Figure 4-1.

Power assurance check (Sheet 1 of 2) Rev. 14

4-7


BHT-412-FM-4

MODEL412

POWER ASSURANCE CHECK- IN-FLIGHT PTST-3D ENGINE

ESTABLISHLEVEL FLIGHT ABOVE 1000 FEET AGL.

HEATER/ECU-OFF.

ENTERCHART AT INDICATEDENGINETORQUE,MOVE UP TO INTERSECTPRESSUREALTITUDE,PROCEEDTO THE RIGHT TO INTERSECTOUTSIDEAIR TEMPERATURE, THEN MOVE UP TO READ VALUE FOR MAXIMUMALLOWABLE ITT.

THROTTLES: TEST ENGINE-FULL OPEN,FRICTIONED.

IF INDICATEDITT EXCEEDS MAX ALLOWABLE,REPEAT CHECKSTABILIZING POWERFOUR MINUTES.

AIRSPEED-100 KIAS (OR VNE, IF LESSI.

OTHERENGINE-DECREASESLOWLYUNTIL TEST ENGINE TORQUEIS WITHINTEST RANGE.DO NOT EXCEED 810°C ITT, 103.1%N1 RPM, OR 73.2% TORQUE.

N11RPM-97%.

REPEATCHECK USINGOTHERENGINE.

IF EITHERENGINEEXCEEDSALLOWABLE ITT AFTER STABILIZINGFOUR MINUTES,PUBLISHEDPERFORMANCE MAY NOT BE ACHIEVABLE.CAUSE SHOULDBE DETERMINED AS SOON AS PRACTICAL.

STABILIZE POWERONE MINUTEIN LEVEL FLIGHT, THEN RECORDPRESSUREALTITUDE,OAT, ENGINETORQUEAND ITT.

MAXIMUM ALLOWABLE ITT-C

40

50

500

550

60

70

B00

650

700

750

800

ENGINE TORQUE-PERCENT (INDICATED)

412EP-FM-4-1-2

Figure 4-1.

4-8

Rev. 14

Power assurance check (Sheet 2 of 2)


BHT-412-FM-4

38

1 80

(11.0)-

36

1 76

1

EXAMPLE IF AMBIENT TEMP IS -15"C AND PRESSURE ALT IS 6.000 FEET. THE DENSITY ALT IS 4,000 FEET AND IS 1 06

1.74

- 1.72 -1.68

1 70

(7.0)-

(6.0)-

(5.0)-

(4.0)-

(3.0)-

TEMPERATURE - °C

412EP-FM-4-2

Figure 4-2.

Density altitude

Rev. 2

4-91


BHT-412-FM-4

NOTE: 1. Pedalcritical wind azimuth-hovering with the relative wind within these azimuth angles can result in inability to maintain heading due to large left pedal requirements for certain wind velocities. 2. Longitudinal cyclic critical wind azimuth aft cyclic may be limited with AFCS longitudinal hardover.


SEENOTE 1


FAA APPROVED

BHT-412-FM-4

T N I R P N O R 0 I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER CEILING

OUT OF GROUND EFFECT

SKID HEIGHT 60 FT. HEATER OFF 0 TO 52°C

TAKEOFF POWER ENGINE RPM 100% GENERATOR 150 AMPS(EA.)

CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

0

14.000 FT HD LIMIT

// ;

0

10

20

OAT -

,

30

C

Z

40

\

AREA A

50

8

|

/

9

10

GROSS WEIGHT -

11

12

LBS x 1000

I I I I I I I I I '

3.4

3.8

4.2

4.6

5.0

I

5.4

GROSS WEIGHT - KG X 1000

412-EP-FM-4-4-1

Figure 4-4. Hover ceiling OGE (Sheet 1 of 12)

SEP 05 000

Rev. 21

411

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER CEILING OUT OF GROUND EFFECT


SKID HEIGHT 60 FT. HEATER OFF -40 TO 0°C

CAUTION: OGE HOVER OPERATION MAY RESULT IN VIOLATION OF H-V LIMITATIONS

14,000 FT HD LIMIT

-40I

_-/ 7~ %i

_7AO/

R:A AA -R

I

-

-40

-30

-20

-10

OAT -

°C

0

8

9

10

11

12

GROSS WEIGHT -

LBS x 1000

I

I

3.4

I

3.8

'

I

4.2

I'

4.6

GROSS WEIGHT -

5.0

I

5.4

KG X 1000

412-EPzFM4z4-2

Figure 4-4. Hover ceiling OGE (Sheet 2)

4-12

Rev. 21

SEP 05 2000

12-EP-FM-

FAA APPROVED

BHT-412-FM-4


TAKEOFF POWER SKID HEIGHT 60 FT. ENGINE RPM 100% HEATER ON GENERATOR 150 AMPS(EA.) 0 TO 20°C CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

14,000 FT HD LIMIT

,

Ht1

AR EA A

0

10

20

OAT - °C

8

9

10

11

12

GROSS WEIGHT - LBS x 1000 I I I I I I I I

3.4

3.8

4.2

4.6

GROSS WEIGHT -

5.0

5.4

KG X 1000

412-EP-FM-4-4-3


SEP 0 5 2000

Rev. 21

4-13

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER CEILING


SKID HEIGHT 60 FT. HEATER ON -40 TO O°C



14,000 FT HD LIMIT

A-C

-

4-14

Rev. 21/

SEP_i05 2000^

___

_

11

1

0

1

-

_

_

_

12

-|

2î^^^z~zz~zmI -

-40-51

0

OAT / z _ // OAT - 'GROSS

7

0

0

__910111

GROSS WEIGHT-LBS x 1000 [\ ZAREAA A WEIGHT - KG X 1000

GROSS WEIGHT - KG X 000w

Figure 4-4.

Hover ceiling OGE (Sheet 4)

42-EP-FM-4-4-4

(J 0

BHT-412-FM-4

FAA APPROVED


SKID HEIGHT 60 FT. WINTERIZATION HEATER ON 0 TO 20°C



14,000 FT HD LIMIT

,-X~ __ z __ _

_ _

0

10

9

8

I

I

I

3.8

I

I

4.2

I

4.6

GROSS WEIGHT -

~

12

11

10

GROSS WEIGHT - LBS x 1000

3.4

0Di~~~

0-

N_

AREA A

20

OAT - °C

_

I

I I

5.0

I

5.4

KG X 1000

412-EP-FM-4-4-5

~Figure 4-4.


SEP 05 2000

Rev 21

4-15

BHT-412-FM-4

FAA APPROVED


SKID HEIGHT 60 FT. WINTERIZATION HEATER ON ° -40 TO 0 C



14,000 FT HD LIMIT

.__/

*^16

e

-40

-30

___I Aev.

/-

-20 OAT -

-10 0 °C

'A^T- °c

I

8k

S^p-4

AREA A

11 12 9 10 8 GROSS WEIGHT - LBS x 1000

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

412-EP-FM-4-4-6

Figure 4-4.

|

4-16

Rev. 21

SEP 0 5 2000


FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F I C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER CEILING OUT OF GROUND EFFECT

MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR 150 AMPS(EA.)



14.000 FT HD LIMIT

-

*

l_ ,:-----

0__-kT

\10

SEP=::0:2000

" _zJ_^

/=^ ^ ^ ---

0

10

20

OAT-

^

^

30 °C

40

50

~Figure 4-4.

4-17

8

9

10

11

12

GROSS WEIGHT - LBS x 1000 3.8

4.2

4.6

GROSS WEIGHT -

0b~~~~

Rev. 21

'~~~~~~~~~~~~~~~~~~ ---

^^:z=: -^-

3.4

~

40

5.0

5.4

KG X 1000

412-EP-FM-4-4-7


SEP 0 52000 Rev. 21

4-17 m

BHT412-FM-4

FAA APPROVED

T N I R P N O R I O T F C 0 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER CEILING OUT OF GROUND EFFECT

MAXIMUM CONTINUOUS POWER SKID HEIGHT 60 FT. ENGINE RPM 100% HEATER OFF GENERATOR 150 AMPS(EA.) -40 TO 0°C CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

OAT - °C

14,000 FT HD LIMIT

as~--^^^_40

4-18

/

Rev. 21

-40

-

-

--

-

1 '-I-'

-3,p'-

SEP 0/^_zzz:iz 5 20 _-'' _ ^ ^ - _ _-/_ _ _ _ _ / /_

-30

-20

-10

OAT -

0

8

9

10

0

_

11

12

GROSS WEIGHT -!C LBS x 1000

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

412-EP-FM-4-4-8


4-18

Rev. 21

SEP 0 5 200~

FAA APPROVED

BHT-412-FM-4

T N . I R P N O R I O T F C U T 0 O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0^--I O HOVER CEILING OUT OF GROUND EFFECT


SKID HEIGHT 60 FT. HEATER ON 0 TO 20°C


14,000 FT HD LIMIT

-'

----

I I

M-

AREA A

0

10

20

8

9

10

GROSS WEIGHT -

OAT - OC

3.4

II

3.8

I -I 4.2

I 4.6

11

12

LBS x 1000

II 5.0

I 5.4


412-EP-FM-4-4-9

Figure 4-4.


SEP 05 200

Rv. 21

4-19

BHT-412-FM-4

FAA APPROVED



SKID HEIGHT 60 FT. HEATER ON ° -40 TO 0 C


14.000 FT HD LIMIT

I

I

I r

I

3.4

I

/-'L_

I I

3.8

4.2

OAT - °C

-3o,20 .

'I

4.6

GROSS WEIGHT -

I

5.0

5.4

KG X 1000

412-EP-FM-4-4-10


4-20 4-20

_ _ /_ SEP_0 5 2D0

Rev. 21 Rev. 21

SFp 0 5 2[[

^

^

^

_

_

FAA APPROVED

BHT-412-FM-4





14.000 FT HD LIMIT

OAT -

AREA A

p0, ^

0

10

C

20

8

OAT - °C

9

10

11

12

GROSS WEIGHT - LBS x 1000 I II I I' I I I i

3.4

3.8

4.2

4.6

GROSS WEIGHT -

5.0

5.4

KG X 1000

412-EP-FM-4-4-11

Figure 4-4.


SEP05 2O. Rev. 21

4-21

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O HOVER CEILING OUT OF GROUND EFFECT

SKID HEIGHT 60 FT. WINTERIZATION HEATER ON -40 TO 0°C



14,000 FT HD LIMIT

' ~_,o/

-40

-30

'Cc

OAT

_

-20 -10 OAT - ,C

0

11 12 9 10 8 GROSS WEIGHT - LBS x 1000 I I I l I I I I I 3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT -

KG X 1000

412-EP-FM-4-4-12


4-22

Rev. 21j

4-22

Rev. 21

SEP _ OJz5 2

_

r'~EP 0 5 2QQQ

_ /fzzrz'~

_

_ :

_

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I _\Y O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TAKEOFF DISTANCE OVER 50 FOOT OBSTACLE

HOVER POWER * 15% TORQUE ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

I^B

OOO-14,000 FT.

/iI

7

-~-

-?

INITIATED FROM 4 FT. SKID HEIGHT VTOCS = 45 KIAS HEATER ON OR OFF

DEN. ALT. LIMIT

i 11/ r _i i /

i---

----

_I.-

---

MAXIMUM

--

GROSS WEIGHT

FOR TAKEOFF

_z--MAX

OAT HEATER ON

*

*I

'--

6 ---0-420 0 20 40 60 400 OAT - °C

//-'

600 800 1000 1200 TAKEOFF DISTANCE - FT.

1400

ENGLISH

412EP-FM-4-5-1

0I

Figure 4-5. Takeoff distance (Sheet 1 of 2)

SEP 05 2000

Rev. 21

4-23

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TAKEOFF DISTANCE OVER 15 METER OBSTACLE

HOVER POWER + 15% TORQUE ENGINE RPM 100% GENERATOR 150 AMPS (EA)

INITIATED FROM 1.2 meter SKID HEIGHT VTOCS= 45 KIAS HEATER ON OR OFF

4267 meter DEN. ALT. LIMIT

-

/ -------k--

--

O--

AT- °C

~MAXIMUM GROSS WEIGHT

TAKEOF

DI

E-FOR TAKEOFF

TMRI~~~

MIN

OAT

-MAX OAT

__

MAX OAT

OT- I;

-60-40-20

OA

/HEATER

0 20 40 60 -

ON/-

100

OHEATERNC

200

300

400

TAKEOFF DISTANCE-

METRIC

412EP-FM-4-5-2

Figure 4-5. Takeoff distance (Sheet 2)

4-24

Rev. 21

SEP b52w

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 7000 LB (3175 kg)

70 KIAS

TAKEOFF POWER

HEATER OFF

ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

1. CUMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN

WITH ALL DOORS OPEN OR REMOVED:

~RATE OF CLIMB - M/SEC

~

i^B-~

0

1 2 3 4 5

20000 III

t'lll

Hll llll l

6 7

tlllImll i

8 9 10 11 12 13 14 15 16 17 18 iihi .. I hliiiimh ml iIl lll'l fl

\\

00000

-4

18000

16000 ol2000 _

,., 14000 L-

____ _ _

_ _tv

\

\_ *^\

_______3°

-: 0 _O__ a~~~~~~~~~~~~~~~~~~~~~~~,,ooo----^\\S^ _________ y.___ 1X __o§~~~~ 0

0

t: *

1000 8000

\3

V,

0b~~~ ~

w

X----

0--h?

~Figure

_______________41 ____________________________

2-EP-FM -4-6-1

4-6. Twin engine rate of climb (Sheet 1 of 24)

SEPO 5 2o'~t

S~P 0 5

Rev. 21

4-25

Re. 21

4-25

BHT-412-FM-4

FAA APPROVED


TAKEOFF POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

70 KIAS HEATER ON

1. CUMB SPEED IS 60 KIAS 2. RATE OF CUMB WILL DECREASE 275 FT/MIN


RATE OF CLIMB -

0

1 2 3 4 5 6

20000l

_

l

liiililiil

l

___

-

16000

M/SEC

9 101112 131415161718 ill

18000

,

7 8

-

-

-_

--

12000

60000 8000 -

y>

400

Rev. 21

J

M

~~~~~43--w

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB - FT/MIN

800

Figure 4-6. Twin engine rate of climb (Sheet 2)

4-26

\.

4-6. Twin engine rate of climb (Sheet 2)

4000igure

0

\

---

SEP 0 5 2;

i

412-EP-FM-4-6-2

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S 0 N G O N I R W O E I F Y V L * o N O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 8000 LB (3629 kg)

TAKEOFF POWER

70 KIAS


HEATER OFF


20000

RATE OF CLIMB - M/SEC 5 6 7 8 9 10 11 12 13 14 1516 17 18

1 2 3 4

0

1. CLMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN

L

18000

''4

----

16000

0

^

0

\_

0-- __ --

uj 12000

>

N

^

S

^

\

^

___40

0

_X

-

8000

Q-

6

LL

---

6000

12001

-

2000 -

----

0

400

------

800

----

0 I00

------

'- 2800 3200 36 4jj-

---

:

I


412-EP-FM-4-6-3


SEP 0 5

Rev.

21

4-27

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 8000 LB (3629 kg)


70 KIAS HEATER ON


0

1 2 3

20000


RATE OF CLIMB - M/SEC 4 5 6 7 8 9 1011 12131415161718 llil

iUll Ll

U

\ \

18000 -_

\

^

-

16000

-5

IIi

0

___

___Iv__

_____

E

ooo 42 ~Re-2 1oo4000-------------,Hflx w 12000

.

-

--

\r

6

1000

30

4000

2000 I-

[

0

400

800


412-EP-FM-4-6-4

Figure 4-6. Twin engine rate of climb (Sheet 4)-EP-FM-4-4

4-28

Rev. 21

SEP 0 5 2000

FAA APPROVED

BHT-412-FM-4


70 KIAS

TAKEOFF POWER


HEATER OFF



RATE OF CLIMB -

0 1 2 3 4 5 6

M/SEC

7 8 9 10 11 12 13 14 15 16 17 18

20000

18000

16000,, w 14000

"

j

a

'

x,

4,

,

S

12000

______

~oooo

x,\_\

-

^^^Lr

-

-

0

1

0

o

8000

-

-

I412-EP-FM-4-5 'i

4000

6000S 20


,,N IIII

400

800

Rev 21

4-29


412-EP-FM-4-6-5


SEP 0 5 2000 Rev. 21

4-29

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N 1 G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 9000 LB (4082 kg)

70 KIAS HEATER ON




0

20000

RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 .,.;.S,

h.

1

]J

6 ~1u

0

18000-

~,4eooo~ ^ ^

^

S

^

ZmJ

P40

w 12000 _

C ,,ooo

Li.

80001

rY

o.

0

vrw_

6000

w

4000 2000

0

400

800



4-30

Rev. 21

SEP 05 2000

412-EP-FM-4-

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L D o-^ N O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB (4536 kg)


70 KIAS HEATER OFF


OF CLIMB - M/SEC ~RATE 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 iii u

~)^B-~

0

20000

-

18000


'- -

6

- 5uIIIIII

16000

w 14000

,

0

12000

_5 "1ooooLiic~~~"

-

CI

O RATE

-----

--

40

-

"-3

I

8000

Liiw~

0

a

…

6000…

\

Q.

4000

\ _

2000\

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FTr/MIN

412-EP-FM-4-6-7


SEP 052000

Rev. 21

4-31

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C 0 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB (4536 kg)


70 KIAS HEATER ON



RATE OF CLIMB - M/SEC 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 20000 I L J ll il'iii L ii wl lnil l ul6ii 18000s

-

-

16000 120161

L I-

La-

61

-

-----DO

--

--

s0 3600 3 2 0 O

14000

A I-

,.12000

%!'

o

0

-3

10000

<

8000

I

Q

2-

---

-

--

6000

4000

2000

-

-

-

-

-

-

-

-

-

-

-

- -

-

-

-

-

- - 3

Ill

0

400

800



4-32

Rev. 21

SEP 0 5 2000

412-EP-FM-4-6-8

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 11,000 LB (4990 kg)

70 KIAS

TAKEOFF POWER

HEATER OFF




20000

RATE OF CLIMB - M/SEC 5 6 7 8 9 10 11 12 13 14 15 16 17 18 ihl luj{l'l l [ l' Illumlliwli'l2id .1'| rllii'' l1u1 Xiiiliiu'ilal | \ 6

1 2 3 4

0

LU

18000

16000

,

14-000

___

12000

0

_

_I-- _

0

AE OF CLIMB - FT

0)8000

6000o.

---

%aJ

-----

----- ^-SEP

0 5200

Rev. 21

4

4000

-

2000

00 0

400

800


412-EP-FM-4-6-9


SEP 0 5 2000 Rev. 21

4-33

BHT-412-FM-4

FAA APPROVED

T N I R P N 0 O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 11,000 LB (4990 kg)

TAKEOFF POWER ENGINE RPM 100%

70 KIAS HEATER ON

GENERATOR 150 AMPS (EA.)


0

20000

1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN

RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 l

ill

i l

l

l

l

Iilll

lllll

ll 1iI

l mll

6I

111 if

1800016000

14000-

-s

:,

"

o

0

^R 10000 12000

-

-3

0

\-

Q:

2-

__

SEP Figure O 2---

6000----------

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FT/MIN

412-EP-FM-4-6-10


4-34

Rev. 21

SEP 0 5 2000,

2-----0

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 11,900 LB (5398 kg)

70 KIAS

TAKEOFF POWER

HEATER OFF




0

20000

l

RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 9 1011 12 131415161718 Ii iiilid iLl llll L Ilil h I il L i ilii

--------

18000

~!

16000

i ,

---

s-'

14000

4t2 EP-FM-4-6-tt 2

----

40

|0

---

8000

1000 4

L.'X a. \ fl V I-1

\

4 000-----

3

--

-L

2

I

0 a ^-

-

2000

o0 0

…

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FT/MIN

412-EP-FM-4-6-11


SEP

05 2900 Rev. 21

4-35

BHT-412-FM-4

FAA APPROVED



70 KIAS HEATER ON




0

20000

im i UI

il

Ill

i ,l llil U

iii

ml l

i

i

ii ul

il

--------

18000 16000

8

'

--

i-

0

J 14000

-i

_

-^__

L D12000

Figu Ti \ en r

1000----

I

4

2000

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FT/MIN


4-36

Rev. 21

SEP 0 5 2009

412-EP-FM-4-6-12

FAA APPROVED

BHT-412-FM-4


MAXIMUM CONTINUOUS POWER

70 KIAS


HEATER OFF


I^Ht~

0



~

1 2

3 4 5

6 7 8 9 10 11 12 13 14 15 16 17 18

6

20000

*

18000

_XXi

16000

----

1 14000

_°

X

412-EP-FM-4M40

i

00 0

w 12

10000

30

8000---…-

I

I

U)

i4

46000

/

a.

2000

0

4000 -------

400

800

0

0H±^--

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB

- FT/MIN

412-EP-FM-4-6-13


SEP 0 5 2D

Rev. 21

4-37

BHT-412-FM-4

FAA APPROVED


MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

70 KIAS HEATER ON


RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

0 1 2

20000


Ii

"ILLLLI

I

'

1

'±d

wPd

6iu

i

18000

)

16000

u . D\_\O

_I

L__]_

_ 14000

£-

-

-

0-4

- 10000 oooo3

<

8000

46

iFa. 6000

I

4000

---------

2000 /

/

Uw

0

400

800

0 1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FT/MIN


4-38

Rev. 21

SEP 05 2000

412-EP-FM-4-6-1 4

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L 0 N O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 8000 LB (3629 kg)


70 KIAS


HEATER OFF



~

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0

20000

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30

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46000

4000--… 0 400

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800

1200 1600 2000_______ 2400 2800 3200 3600 1l' _

RATE OF CLIMB - FT/MIN

412-EP-FM-4-6-15


SEP 0 5 2000

Rev. 21

4-39

BHT-412-FM-4

FAA APPROVED

T N I R P N 0 O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 8000 LB (3629 kg)


70 KIAS HEATER ON



RATE OF CLIMB - M/SEC 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

20000 18000

Il

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8000

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4000

4000

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0

0

400

800

1200

1600 2000 2400 2800

RATE OF CLIMB -


4-40

Rev. 21

SEP

0 5 2000

3200 3600

FT/MIN

412-EP-FM-4-6-16

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C -. U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 9000 LB (4082 kg)

70 KIAS HEATER OFF



0

RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 i

1 2

20000


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18000

16000

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400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB

-

FT/MIN

412-EP-FM-4-6-17


SEP 052000

Rev. 21

4-41

FAA APPROVED

BHT-412-FM-4


70 KIAS HEATER ON



0

1 2 3


RATE OF CLIMB - M/SEC 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

6

20000

"'

18000

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2000

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1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

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412-EP-FM-4-6-18


4-42

Rev. 21

SEP 05 2000

BHT-412-FM-4

FAA APPROVED



70 KIAS HEATER OFF



OF CLIMB - M/SEC I::~~RATE 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 tlliii lliiiiiiiillldImlllllmliI i6nnr l l mlmImmllmf mm lmmllImll

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412-EP-FM-4-6-19

Figure 4-6.

Twin engine rate of climb (Sheet 19)

SEP 05 2000Rev. 21

4-43

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C 0 U T O OD N R E P E E R R C S N I -/ G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB (4536 kg)


70 KIAS HEATER ON


1. CLIMB SPEED IS60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FTIMIN

RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

0

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20000

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l

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400

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1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

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4-44

Rev. 21

4

SEP 0 5 2000

412-EP-FM-4-6-20

(

FAA APPROVED

BHT-412.FM-4



70 KIAS


HEATER OFF


0

20000


RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 1011 12 131415161718

1 2

Ii

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SEP 0

0O

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

0

FT/MIN

412-EP-FM-4-6-21


SEP O5 20O0

Rev. 21

4-45

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE RATE OF CUMB GROSS WEIGHT 11,000 LB (4990 kg)


70 KIAS HEATER ON


0

1 2

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1200

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2800 3200

3600

FT/MIN

412-EP-FM-4-6-22

Figure 4-6.

4-46

Rev. 21

Sp

0 5 2000


41

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P 0 E E R R C S N G O N o I R W O E I F V NLY O TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 11,900 LB (5398 kg)


70 KIAS


HEATER OFF



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32600

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Re. 21 Re.21r

4-47 447

BHT-412-FM-4

FAA APPROVED


70 KIAS HEATER ON



0

20000

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24)

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1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

4-48

5_--

,

10000

0.

4

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FT/MIN


2000

412EP-FM-4-6-24

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C zL U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 7000 LB (3175 kg)


70 KIAS

ENGINE RPM 97% GENERATOR 150 AMPS

HEATER OFF INOPERATIVE ENGINE SECURED


-4-3-2-1 20000 iill

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0


RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 i li

18000

9

10 11 12 -6

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16000

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400

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1200

-

1600

2000

2400

FT/MIN

Figure 4-7. Single engine rate of climb (Sheet 1 of 6)

SEP 05 2000

Rev. 21

4-49

BHT-412-FM-4

FAA APPROVED

T N I R P N 0 O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CUMB GROSS WEIGHT 8000 LB (3629 kg)

MAXIMUM CONTINUOUS POWER ENGINE RPM 97% GENERATOR 150 AMPS

70 KIAS HEATER OFF INOPERATIVE ENGINE SECURED



RATE OF CLIMB - M/SEC -6-5-4-3-2-1 0 1 2 3 4 5 6

20000

7

8

9

10 1

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18000

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4-50

Rev. 21 2000

-1200

SEP 0 5 zao -800

-400

0

400

800

1200

1600

2000


412-EP-FM-4-7-2

Figure 4-7. Single engine rate of climb (Sheet 2)

4-50

Rev. 21

SEP 05 2000

4

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 9000 LB (4082 kg)


70 KIAS




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1200

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SEP 052000

Rev. 21

4-51

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CUMB GROSS WEIGHT 10,000 LB (4536 kg)


70 KIAS




1. CLIMB SPEED IS 60 KIAS 2. RATE OF CUMB WILL DECREASE 275 FTIMIN

RATE OF CLIMB - M/SEC -10-9 -8-7-6-5-4-3-2 -1 0 1 2

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-2000-1600 -1200 -800

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RATE OF CUMB -

0

400

800

1200

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412-EP-FM-4-7-4

Figure 4-7.

4-52

Rev. 21

Single engine rate of climb (Sheet 4)

$SEP052000

FAA APPROVED

BHT-412-FM-4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 11,000 LB (4990 kg)





RATE OF CLIMB - M/SEC -10-9 -8-7-6-5-4 -3-2-1 0 1 2

3

4

5

6

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SEP05200 SEP 05 2000

Rev. Rev. 21 21

453 4-53

BHT-412-FM-4

FAA APPROVED

T N I R P N 0 O R I O T F C U T O OD N R E P E E R R C S N G T O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 11,900 LB (5398 kg)





RATE OF CLIMB - M/SEC -10-9 -8-7 -6 -5 -4-3-2-1 0 1 2 20000 i

3

4

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5 i

6

18000

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412-EP-FM-4-7-6


4-54

Rev. 21

SEP 0 5 2000

412-EP-F-4-7-

BHT-412-FM-4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O PILOT & COPILOT AIRSPEED SYSTEM CALIBRATION

CLIMB. LEVEL FLIGHT, AUTOROTATION

SKID GEAR 160---

KIAS -

ERROR = KCAS

140

LEVEL FLIGHT

U3 -

iO

120

40 -…/

I l

l

80 0

20

i

40

I

,

AUTOROTATION-.I

60

80

100

120

140

160

20

Figure4-8. Airspeed calibration

SEP 052000

Rev.21

4-55

BHT-412-FM-4

FAA APPROVED

T N I R P N O I 0OR T F C U T D O MIN.1 / O N R E P E E 7 R R C S N G O N I _ - R _… iii_ W O E I F V NLY O SINGLE ENGINE LANDING DISTANCE OVER 50 FT. (15 M) OBSTACLE

2/2 MINUTE OEI POWER AS REQUIRED ENGINE RPM 97% GENERATOR 150 AMPS

45 KIAS RATE OF DESCENT 500 FT/M HARD SURFACED RUNWAY INOPERATIVE ENGINE SECURED

14,000 FT. DEN. ALT. LIMIT

^

I -IIIg -I--

1

MIN.-

40 20

OAT

20 40 60 200

oA T

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7 I/. /

™MAX.

400

600

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- OC

-40 -20 0

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20 40 60 200

OAT - °C

' 50

/~ ^

1000

600

' 150

I 200

800

I 250

1000 I 300

LANDING DISTANCE

Figure 4-9. Single engine landing distance

4-56

Rev. 21

SEP 0 5 2000

1200

FEET

LANDING DISTANCE ALL GROSS WEIGHTS____

400

I 100

800--

1200

I 350

FEET

I' ' 400 METERS

W

412EP-FM-4-9

0

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

BHT-412-FM-4

TABLE OF CONTENTS

Subject

Paragraph

Page Number

INTRODUCTION ............................ ..................... EMPTY WEIGHT CENTER OF GRAVITY ........................ GROSS WEIGHT CENTER OF GRAVITY........................ DOORS OPEN OR REMOVED.................................. COCKPIT AND CABIN LOADING ............................... BAGGAGE COMPARTMENT LOADING ......................... FUEL LOADING ................................................. SAMPLE LOADING PROBLEM ................................. SAMPLE LOADING PROBLEM (ENGLISH)................. SAMPLE LOADING PROBLEM (METRIC) .................. REQUIRED EQUIPMENT LIST..................................

5-1........... 5-2 .......... 5-3 .......... 5-4........... 5-5........... 5-6........... 5-7 .......... 5-8 .......... 5-8-A ........ 5-8-B ........ 5-9 ..........

5-3 5-3 5-3 5-3 5-3 5-4 5-4 5-4 5-4 5-7 5-9

LIST OF FIGURES

Title

Figure Number

Helicopter station diagram ..................................... Internal fuel tank station location ..............................

5-1........... 5-2...........

Page Number 5-11 5-12

LIST OF TABLES

Title

Table Number

Door Weights and Moments (English) ......................... Door Weights and Moments (Metric)........................... Pilot and passenger table of moments (English).............. Pilot and passenger table of moments (Metric) ............... Internal cargo loading table (English) ......................... Internal cargo loading table (Metric) ........................... Baggage loading table (English) ............................... Baggage loading table (Metric) ................................ Fuel loading table (English) .................................... Fuel loading table (Metric) ...................................... Fuel loading table - lateral (English) ......................... Fuel loading table - lateral (Metric)...........................

5-1........... 5-1M ......... 5-2........... 5-2M ......... 5-3........... 5-3M ......... 5-4........... 5-4M ......... 5-5........... 5-5M ......... 5-6........... 5-6M .........

Page Number 5-13 5-13 5-14 5-14 5-15 5-16 5-17 5-18 5-19 5-20 5-21 5-22

5-1/5-2

T N I * Section5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

5-1. INTRODUCTION

This section provides information regarding helicopter center of gravity and cockpit and cabin loading. Loading tables for pilot, passengers, cargo, and fuel are provided. A sample loading problem is provided to aid in flight planning.

5-2. EMPTYWEIGHTCENTER OF GRAVITY

The empty weight consists of the basic helicopter with required equipment, optional equipment kits, transmission and gearbox oils, hydraulic fluid, unusable fuel, undrainable engine oil, and fixed ballast. The empty weight center of gravity shall be adjusted within the limits of the applicable Weight Empty Center of Gravity chart in the Maintenance Manual.

5-3. GROSSWEIGHTCENTER OF GRAVITY

It shall be the pilots responsibility to ensure that the helicopter is properly loaded so that the entire flight is conducted within the limits of the Gross Weight Center of Gravity chart in the Limitations section of the Flight Manual. The gross weight center of gravity may be calculated from the helicopter Actual Weight Record (historical records) and the Loading Tables in this section or in appropriate Flight Manual Supplements to assure safe loading.

BHT-412-FM-4

Figure 5-1 presents fuselage station and buttock line data to aid in weight and balance computations.

5-4. DOORSOPEN OR REMOVED

Opening or removing doors results in center of gravity changes. Door configuration shall be symmetrical for both sides of the fuselage. Tables 5-1 and 5-1M list weight and moment adjustments which should be made in determining the gross weight and CG when a pair of doors are opened or removed.

5-5. COCKPITAND CABIN LOADING

A minimum crew weight of 170 pounds (77.1 kilograms) in the cockpit is required. Except for the two aft passenger seats, crew and passengers may be loaded In any sequence without exceeding the gross weight center of gravity limits approved for flight.

Refer to Tables 5-2 and 5-2M for personnel weights and moments in both English and metric units. NOTE

The two aft outboard facing seats should not be occupied unless at least four passengers are seated in the forward or aft facing seats.

The cabin deck cargo loading limit is 100 pounds per square foot (4.9 kg/100 sq cm). 5-3


WARNING

HELICOPTER CENTER OF GRAVITY SHALL BE COMPUTED FOR ALL CARGO BAGGAGE/ CONFIGURATIONS BEFORE FLIGHT.

Refer to Tables 5-3 and 5-3M for internal cargo weight and moment data.

5-6. BAGGAGE COMPARTMENTLOADING

The baggage compartment is accessible from the right side of the tailboom and contains approximately 25 cubic feet (0.8 cubic meters) of space. The baggage compartment has a load limit of 400 pounds (181 kilograms), not to exceed 100 pounds per square foot (4.9 kg/100sq cm). These are structural limitations only and do not infer that CG will remain within approved limits. When weight is loaded compartment, into the baggage indiscriminate crew, passenger and fuel loading can no longer be assumed, and the pilot must compute gross weight CG to assure loading within approved limits. Loading of the baggage compartment should be from front to rear. The load shall be secured to tiedown fittings If shifting of the load in flight could result in structural damage to the baggage compartment or in gross weight center of gravity limits being exceeded. The CG shall be computed with the load In the most adverse position.

5-4

MANUFACTURER'SDATA

Refer to tables 5-4 and 5-4M for baggage weights and moments.

5-7. FUEL LOADING

Due to the fuel flow sequencing between the tanks, the fuel loading CG will vary between fuselage station 139.4 and 160.9. The maximum aft CG will occur at approximately 952 pounds for Jet A, A-1. JP-5 and approximately 910 pounds for Jet B, JP-4. The maximum forward CG will occur at 397 pounds for Jet A, A-1, JP-5 and at 379 pounds for Jet B, JP-4. With normal crew and passenger loading, gross weight CG will remain within limits at any fuel quantity. Refer to tables 5-5, 5-5M and 5-6, 5-6M for fuel weights and moments Figure 5-2 depicts fuel tank location by station number.

5-8. SAMPLE LOADING PROBLEM 5-8-A. SAMPLELOADING PROBLEM(ENGLISH)

The helicopter is chartered to transport nine passengers and 180 pounds of baggage on a trip that will require approximately 260 U.S. gallons of fuel one way. The helicopter will be refueled and the 190-pound pilot will return alone. Determine extreme CG conditions for both flights.


BHT-412-FM-4

OUTBOUNDFLIGHT

LATERAL

LONGITUDINAL

WEIGHT

Weight Empty + Oil

+ Pilot

+ Passengers, (5 man seat) + Passengers, (4 man seat) + Baggage

7000

MOMENT

CG

MOMENT

CG

143.0

1001000

+0.2

+1400

0 +22.0 0 0 0

0 +4180 0 0 0

4146 8930 99450 59160 46980

25 190 850 680 180

Basic Operating Weight + Payload

8925

136.7

1219666

+0.6

+5580

+ Takeoff Fuel (320 U.S. Gallons)

2080

150.7

313456

-0.4

-811

11005

139.3

1533122

+0.4

+4769

8925

136.7

1219666

+0.6

+5580

379

139.9

53022

Most Forward Condition

9304

136.8

1272688

+0.6

+5580


8925

136.7

1219666

+0.6

+5580

390

141.1

55029

0

0

9315

136.8

1274695

+0.6

Takeoff Condition

Basic Operating Weight + Payload + Critical Forward Fuel (58.3 U.S. Gallons)

+ Landing Fuel (60 U.S. Gallons)

Landing Condition

0

0

+5580

5-5


BHT-412-FM-4

RETURNFLIGHT

LATERAL

LONGITUDINAL

WEIGHT

Weight Empty + Oil + Pilot

7000

MOMENT

CG

MOMENT

CG

143.0

1001000

+0.2

+1400

4146 8930

+22.0

+4180

25 190

Basic Operating Weight

7215

140.6

1014076

+0.8

+5580

+ Takeoff Fuel (320 U.S. Gallons)

2080

150.7

313456

-0.4

-811

Takeoff Condition

9295

142.8

1327532

+0.5

+4769


7215

140.6

1014076

+0.8

+5580

+ Critical Forward Fuel (58.3 U.S. Gallons)

379

139.9

53022


7594

140.5

1067098

+0.8

+5580


7215

140.6

1014076

+0.8

+5580

390

141.1

55029

7605

140.6

1069105

+ Landing Fuel (60 U.S. Gallons)

Landing Condition

5-6

0

0

+0.7

0

0

+ 5580


BHT-412-FM-4

5-8-B. SAMPLE LOADING

a trip that will require approximately

The helicopter is chartered to transport nine passengers and 80 kg of baggage for

alone. Determine for both flights.

extreme

960

CG conditions

OUTBOUND FLIGHT

LATERAL

LONGITUDINAL

Weight Empty

+Oil + Pilot + Passengers, (5 man seat) + Passengers, (4 man seat) + Baggage

MOMENT kgemm 100

WEIGHT (kg)

CG (mm)

MOMENT kgemm 100

CG (mm)

3175.1

3632

115319.6

+5

485.3 1074.6 11145.0 6630.0 5296.0

0 +559 0 0 0

0 +503 0 0 0

11.3 90.0 375.0 300.0 80.0

+159

4031.4

3472

139950.5

+16

+ 662

934.7

3822

35724.2

-10

-93

Takeoff Condition

4966.1

3537

175674.7

+11

+ 569


4031.4

3472

139950.5

+16

+ 662

171.9

3553

6107.6

0

0


4203.3

3475

146058.1

+16

+ 662


4031.4

3472

139950.5

+16

+662

186.9

3635

6793.8

-1

-2

4218.3

3479

146744.3

+16

+660

Basic Operating Weight + Payload + Takeoff Fuel (1200 Liters)

+ Forward Fuel(220.7 liters)

+ Landing Fuel (240 Liters)

Landing Condition

5-7


BHT-412-FM-4

RETURN FLIGHT

LATERAL

LONGITUDINAL

WEIGHT (kg)

Weight Empty + Oil + Pilot

3175.1

CG (mm)

3632

11.3 90.0

MOMENT kgomm 100

CG (mm)

MOMENT kg*mm 100

115319.6

+5

+159

485.3 1074.6

0 +559

0 +503

3276.4

3567

116879.5

+20

+662

934.7

3822

35724.2

-10

-93

Takeoff Condition

4211.1

3624

152603.7

+14

+569


3276.4

3567

116879.5

+20

+662

171.9

3553

6107.6

0


3448.3

3567

122987.1

+19

+662


3276.4

3567

116879.5

+20

+662

186.9

3635

6793.8

-1

-2

3463.3

3571

123673.3

+19

+660


+ Takeoff Fuel (1200 Liters)

+ Forward Fuel (220.7 liters)

+ Landing Fuel (240 Liters)

Landing Condition

0

MANUFACTURER'S DATA

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S

BHT-412-FM-4

DATA

5-9. REQUIREDEQUIPMENT

LIST

listed for each item were included in calculating the empty weight and center of gravity.

A list of basic equipment required for flight is presented. Weights and arms

REQUIRED EQUIPMENT

Indicators Attitude Airspeed Vertical Speed Altimeter Triple Tachometer XMSNOil Press & Temp Gearbox Oil Press & Temp Engine Oil Press & Temp (2) Fuel Pressure (2) Gas Producer Tachometer (2) Turbine Inlet Temperature (2) Hydraulic Oil Press & Temp (2) Dual Torque Pressure Fuel Quantity Standby Compass Horizontal Situation Clock Free Air Temperature Dual DC Ammeter Dual AC/DC Voltmeter (2) Fire Warning Engine No. 1 Engine No. 2 Baggage Compartment Low Fuel WRN - Master Master Caution Pnl Starter-Generator, L.H. Starter-Generator, R.H. Battery Starter Toggle Switch Fuel Igniter Switch (2) Anticollision Light, Upper Anticollision Light, Lower Landing Light Searchlight Position Lights Forward Lower (2) Forward Upper (2) Aft (2) Circuit Breaker Panels (2) Nonessential Bus Switch

UNIT WEIGHT LB/kg

LONGITUDINAL ARM IN/mm

LATERAL ARM (-LEFT, +RIGHT) IN/mm

7.5/3.4 1.0/0.5 1.8/0.8 1.8/0.8 2.5/1.1 0.7/0.3 1.0/0.5 1.0/0.5 ea. 0.6/0.3 ea. 1.0/0.5 ea. 0.8/0.4 ea. 0.4/0.2 ea. 2.6/1.2 1.0/0.5 0.8/0.4 6.5/2.9 0.5/0.2 0.2/0.1 1.0/0.5 1.0/0.5 ea.

21.0/533 22.0/559 23.2/589 22.0/559 23.2/589 25.6/650 25.6/650 24.2/615 24.8/630 23.0/584 23.6/599 23.0/584 23.2/589 23.6/599 38.0/965 22.5/572 25.8/655 40.0/1016 24.0/610 24.0/610

+17.0/+432 +12.5/+318 +21.3/+541 +21.3/+541 +9.0/+229 +6.0/+152 +3.0/+76 +4.7/+119 +4.7/+119 +4.7/+119 +4.7/+119 0/0 +12.5/+318 +1.3/+33 +17.6/+447 +17.0/+432 +26.2/+665 +19.0/+483 -1.3/-41 0/0

0.4/0.2 0.4/0.2 0.3/0.1 6.5/2.9

24.0/610 24.0/610 22.0/559 24.8/630

-1.7/-43 +4.7/+119 -1.8/-46 -2.5/-64

30.0/13.6 30.0/13.6 74.5/33.8 0.4/0.2 0.2/0.1 ea. 1.5/0.7 1.5/0.7 8.0/3.6 5.5/2.5

159.0/4039 159.0/4039 -6.0/-152 38.0/965 33.0/838 169.0/4293 65.4/1661 85.3/2167 50.0/1270

-8.0/-203 +18.0/+457 +8.0/+203 +10.0/+254 +3.0/+76 0/0 0/0 -4.8/-104 +17.6/+447

0.3/0.1 ea. 0.3/0.1 ea. 0.3/0.1 ea. 9.7/4.4 ea. 0.1/NEG.

66.0/1676 109.0/2769 432.0/10973 55.4/1407 41.3/1049

0/0 0/0 0/0 0/0 +3.7/+94

5-9


BHT-412-FM-4

UNIT WEIGHT LB/kg

LONGITUDINAL ARM IN/mm

LATERAL ARM (-LEFT, +RIGHT) IN/mm

R.H. Ovhd Cont Pnl

3.2/1.5

41.3/1049

+3.7/+94

L.H. Ovhd Cont Pnl

2.6/1.2

41.3/1049

-3.7/-94

0.1/NEG. ea.

143.0/3632

0/0

3.8/1.7

10.1/257

-7.8/-198

Control

0.5/0.2

39.0/991

-4.3/-109

Antenna

1.0/0.5

211.1/5362

0/0

Blade and Arm (2)

1.6/0.7 ea.

34.5/876

0/0

Motor (2)

4.2/1.9 ea.

41.0/1041

0/0

REQUIRED EQUIPMENT

Essential Bus

Low Fuel Wrn XMTR (2) VHF No. 1 Radio Transceiver

and Mount

Windshield Wiper -

Fire Extinguisher, L.H

Hand Type,

3.0/1.4

67.8/1722

-35.0/-889

Fire Extinguisher, R.H.

Hand Type,

3.0/1.4

53.5/1359

+34.5/+876

1.7/0.8

Flight Manual

Seat with Restraint -

Pilot

45.4/20.6

Seat with Restraint -

Copilot

45.4/20.6

Seat with restraint -

Copilot*

-/-

54.4/1382

+22.0/+559 -22.0/-559

45.4/20.6

54.4/1382 54.4/1382

-21.2*/-538.5*

Map and Data Case

1.9/0.9

59.0/1499

0/0

First Aid Kit

5.1/2.3

45.5/1156

-7.2/-183

*S/N 36234 helicopters.

5-10

and

Rev. 17

sub

or

modified

BHT-412-FM-4

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O Figure 5-1. Helicopter station diagram

05 NOV 2002

Rev. 24

5-11


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

BHT-412-FM-4

Table 5-1. Door Weights and Moments (English)

DOOR CONFIGURATION

Both crew doors removed Both hinged panels removed Both sliding doors removed Both sliding doors full open

WEIGHT CHANGE

ARM

(POUNDS)

(INCHES)

MOMENT CHANGE (IN-LB)

-39.0

46.2

-1802

-20.4

85.0

-1734

-90.4

130.0

-11,752

0

202.0

+6509

Table 5-1M. Door Weights and Moments (Metric) WEIGHT CHANGE

ARM

DOOR CONFIGURATION

(kg)

(mm)

MOMENT CHANGE (kg-mm/100)

Both crew doors removed Both hinged panels removed Both sliding doors removed Both sliding doors full open

-17.7

1173

-207.6

-9.25

2159

-199.8

-41.0

3302

-1354.0

0

5131

+749.9

5-13


BHT-412-FM-4

Table 5-2.

Pilot and passenger table of moments (English)

PILOT AND PASSENGER TABLE OF MOMENTS (IN-LB)

Weight (Pounds) 100 110 120 130 140 150 160 170 180 190 200 210 220

Pilot and Copilot* F.S. 47 4700 5170 5640 6110 6580 7050 7520 7990 8460 8930 9400 9870 10340

Passenger (4-Man Seat Facing Aft) F.S. 87 8700 9570 10440 11310 12180 13050 13920 14790 15660 16530 17400 18270 19140

Passenger (5-Man Seat Facing Fwd) F.S. 117 11700 12870 14040 15210 16380 17550 18720 19890 21060 22230 23400 24570 25740

PASSENGER FACING OUTBOARD Fwd Seat Aft Seat F.S. 156 F.S. 139 13900 15600 15290 17160 16680 18720 18070 20280 19460 21840 20850 23400 22240 24960 23630 26520 25020 28080 26410 29640 31200 27800 29190 32760 30580 34320

*Left Forward Seat

(TABLE

Table 5-2M.

I.D. 910670)

Pilot and passenger table of moments (Metric)

PILOT AND PASSENGER TABLE OF MOMENTS (kg * mm) 100

Weight (K.G.) 50 55 60 65 70 75 77.1 80 85 90 95 100 105 110

Pilot and Copilot* 1194 mm 597.0 656.7 716.4 776.1 835.8 895.5 920.6 955.2 1014.9 1074.6 1134.3 1194.0 1253.7 1313.4

Passenger (4-Man Seat Facing Aft) 2210 mm 1105.0 1215.5 1326.0 1436.5 1547.0 1657.5 1703.9 1768.0 1878.5 1989.0 2099.5 2210.0 2320.5 2431.0

Passenger (5-Man Seat Facing Fwd) 2972 mm 1486.0 1634.6 1783.2 1931.8 2080.4 2229.0 2291.4 2377.6 2526.2 2674.8 2823.4 2972.0 3120.6 3269.2

PASSENGER FACING OUTBOARD Fwd Seat Aft Seat 3531 mm 3962 mm 1981.0 1765.5 1942.0 2179.1 2118.6 2377.2 2295.1 2575.3 2471.7 2773.4 2648.2 2971.5 2722.4 3054.7 2824.8 3169.6 3367.7 3001.3 3177.9 3565.8 3354.4 3763.9 3531.0 3962.0 3707.5 4160.1 3884.1 4358.2

*Left Forward Seat (TABLE

I.D. 910669)


BHT-412-FM-4

Table 5-3.

Internal cargo

loading table (English)

INTERNAL CARGO LOADING TABLE (ENGLISH) MOMENTS IN-LB

Weight Pounds

F.S. 75.0

F.S. 90.0

50 100 150 200 250

3750 7500 11250 15000 18750

4500 9000 13500 18000 22500

5250 10500 15750 21000 26250

6000 12000 18000 24000 30000

6750 13500 20250 27000 33750

7500 15000 22500 30000 37500

300 350 400 450 500

22500 26250 30000 33750 37500

27000 31500 36000 40500 45000

31500 36750 42000 47250 52500

36000 42000 48000 54000 60000

40500 47250 54000 60750 67500

45000 52500 60000 67500 75000

550 600 650 700 750

41250 45000 48750 52500 56250

49500 54000 58500 63000 67500

57750 63000 68250 73500 78750

66000 72000 78000 84000 90000

74250 81000 87750 94500 101250

82500 90000 97500 105000 112500

800 850 900 950 1000

60000 63750 67500 71250 75000

72000 76500 81000 85500 90000

84000 89250 94500 99750 105000

96000 102000 108000 114000 120000

108000 114750 121500 128250 135000

120000 127500 135000 142500 150000

1050 1100 1150 1200 1250

78750 82500 86250 90000 93750

94500 99000 103500 108000 112500

110250 115500 120750 126000 131250

126000 132000 138000 144000 150000

141750 148500 155250 162000 168750

157500 165000 172500 180000 187500

1300 1350 1400 1450 1500

97500 101250 105000 108750 112500

117000 121500 126000 130500 135000

136500 141750 147000 152250 157500

156000 162000 168000 174000 180000

175500 182250 189000 195750 202500

195000 202500 210000 217500 225000

F.S. 105.0

F.S. 120.0

F.S. 135.0

F.S. 150.0

(TABLE

I.D. 910668)

5-15


BHT-412-FM-4

Table 5-3M.

Internal cargo loading table (Metric)

INTERNAL CARGO LOADING TABLE (METRIC) MOMENTS (kg * mm/100)

Weight (kg) 25 50 75 100

1900 mm 457.0 950.0 1425.0 1900.0

2300 mm 575.0 1150.0 1725.0 2300.0

2700 mm 675.0 1350.0 2025.0 2700.0

3100 mm 775.0 1550.0 2325.0 3100.0

3400 mm 850.0 1700.0 2550.0 3400.0

3800 mm 950.0 1900.0 2850.0 3800.0

125 150 175 200

2375.0 2850.0 3325.0 3800.0

2875.0 3450.0 4025.0 4600.0

3375.0 4050.0 4725.0 5400.0

3875.0 4650.0 5425.0 6200.0

4250.0 5100.0 5950.0 6800.0

4750.0 5700.0 6650.0 7600.0

225 250 275 300

4275.0 4750.0 5225.0 5700.0

5175.0 5750.0 6325.0 6900.0

6075.0 6750.0 7425.0 8100.0

6975.0 7750.0 8525.0 9300.0

7650.0 8500.0 9350.0 10200.0

8550.0 9500.0 10450.0 11400.0

325 350 375 400

6175.0 6650.0 7125.0 7600.0

7475.0 8050.0 8625.0 9200.0

8775.0 9450.0 10125.0 10800.0

10075.0 10850.0 11625.0 12400.0

11050.0 11900.0 12750.0 13600.0

12350.0 13300.0 14250.0 15200.0

425 450 475 500

8075.0 8550.0 9025.0 9500.0

9775.0 10350.0 10925.0 11500.0

11475.0 12150.0 12825.0 13500.0

13175.0 13950.0 14725.0 15500.0

14450.0 15300.0 16150.0 17000.0

16150.0 17100.0 18050.0 19000.0

525 550 575 600

9975.0 10450.0 10925.0 11400.0

12075.0 12650.0 13225.0 13800.0

14175.0 14850.0 15525.0 16200.0

16275.0 17050.0 17825.0 18600.0

17850.0 18700.0 19550.0 20400.0

19950.0 20900.0 21850.0 22800.0

625 650 675 700

11875.0 12350.0 12825.0 13300.0

14375.0 14950.0 15525.0 16100.0

16875.0 17550.0 18225.0 18900.0

19375.0 20150.0 20925.0 21700.0

21250.0 22100.0 22950.0 23800.0

23750.0 24700.0 25650.0 26600.0

(TABLE

I.D. 910667)


BHT-412-FM-4

Table 5-4. Baggage loading table (English) BAGGAGE COMPARTMENT LOADING TABLE (ENGLISH)

NOTE: LOAD BAGGAGE AS FAR FORWARD AS POSSIBLE.

Weight

Approximate CG

Moment

20 40 60 80 100

245 247 249 251 253

4900 9880 14940 20080 25300

120 140 160 180 200

255 257 259 261 263

30600 35980 41440 46980 52600

220 240 260 280 300

265 267 269 271 273

58300 64080 69940 75880 81900

320 340 360 380 400

275 277 279 281 283

88000 94180 100440 106780 113200

(LB)

(Fuselage Sta. -

Inches)

(IN-LB)

(TABLE

I.D. 910666)

5-17


BHT-412-FM-4

Table 5-4M. Baggage loading table (Metric) BAGGAGE COMPARTMENT LOADING TABLE (METRIC)

NOTE: LOAD BAGGAGE AS FAR FORWARD AS POSSIBLE.

Weight (kg)

10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 181.4

Moment kg * mm

Approximate CG

(Fuselage

Sta. -

100

mm)

622.3 1256.0 1900.8 2557.2 3225.0 3903.6 4594.1 5296.0 6009.3 6733.0 7469.0 8216.4 8973.9 9744.0 10525.5 11316.8 12121.0 12936.6 13051.7

6223 6280 6336 6393 6450 6506 6563 6620 6677 6733 6790 6847 6903 6960 7017 7073 7130 7187 7195

(TABLE

I.D.

910665)


BHT-412-FM-4

Table 5-5.

Fuel loading table (English)

USABLE FUEL LOADING TABLE (ENGLISH)

Jet B, JP-4 (6.5 Lbs/Gal)

U.S. Gal. 10 20 30 40 50 *58.3 60 70 80 90 100 110 120 130 140 150 160 170 173.1 180 190 200 210 220 230 240 **243.1 250 260 270 280 290 295.1 300 310 320 330 ***330.5

Weight (Lb) 65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1125 1170 1235 1300 1365 1430 1495 1560 1580 1625 1690 1755 1820 1885 1918 1950 2015 2080 2145 2148

CG In. 139.4 139.6 139.8 139.9 139.9 139.9 141.1 146.0 149.8 152.7 155.0 156.8 158.3 159.7 160.9 156.4 152.4 149.0 148.0 149.0 150.4 151.6 152.8 153.9 154.7 155.7 155.9 154.8 153.2 151.8 150.4 149.1 148.6 149.0 149.9 150.7 151.4 151.5

Jet A, A-1, JP-5 (6.8 Lbs/Gal)

Moment In.-Lb. 9061 18148 27261 36374 45468 53022 55029 66430 77896 89330 100750 112112 123474 134947 146419 152490 158496 164645 166500 174330 185744 197080 208572 220077 231277 242892 246322 251550 258908 266409 273728 281054 285015 290550 302049 313456 324753 325422

U.S. Gal. 10 20 30 40 50 *58.3 60 70 80 90 100 110 120 130 140.0 150 160 170 173.1 180 190 200 210 220 230 240 **243.1 250 260 270 280 290 295.1 300 310 320 330 ***330.5

Weight (Lb) 68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1088 1156 1177 1224 1292 1360 1428 1496 1564 1632 1653 1700 1768 1836 1904 1972 2007 2040 2108 2176 2244 2247

CG In. 139.4 139.6 139.8 139.9 139.9 139.9 141.1 146.0 149.8 152.7 155.0 156.8 158.3 159.7 160.9 156.4 152.4 149.1 148.0 149.0 150.4 151.6 152.8 153.9 154.7 155.7 155.9 154.8 153.2 151.3 150.4 149.1 148.6 149.0 149.9 150.7 151.4 151.5

Moment In.-Lb. 9479 18986 28519 38053 47566 55540 57569 69496 81491 93452 105400 117286 129173 141175 153177 159528 165811 172244 174196 182376 194317 206176 218198 230234 241951 254102 257703 263160 270858 278705 286362 294025 298240 303960 315989 327923 339742 340421

*Most critical fuel amount for most forward C.G. condition. **Most critical fuel amount for most aft C.G. condition at weight empties up to 6750 pounds. ***Most critical fuel amount for most aft C.G. condition at weight empties at 6750 pounds or greater. Weights given are nominal weights at 15°C. NOTE This table is invalid with auxiliary fuel tank(s) installed. (TABLE

I.D.

910664)


BHT-412-FM-4

Table 5-5M.

Fuel loading table (Metric)

USABLE FUEL LOADING TABLE (METRIC)

Jet A, A-1, JP-5 (.815 kg/I)

Jet B, JP-4 (.779 kg/I)

Liters

40 80 120 160 200 *220.7 240 280 320 360 400 440 480 520 530 560 600 640 *655.3 680 720 760 800 840 880 **920 960 1000 1040 1080 *1117.1 1120 1160 1200 1240 ***1251.1

Weight (kg)

CG (mm)

Moment kg-mm

31.2 62.3 93.5 124.6 155.8 171.9 186.9 218.1 249.3 280.4 311.6 342.7 373.9 405.0 412.8 436.2 467.3 498.5 510.4 529.6 560.8 592.0 623.1 654.2 685.4 716.7 747.7 778.9 810.0 841.2 870.1 872.3 903.5 934.7 965.8 974.3

3541 3547 3551 3552 3552 3553 3635 3754 3840 3910 3965 4012 4046 4079 4086 3993 3886 3785 3758 3785 3820 3853 3886 3913 3938 3961 3917 3901 3839 3804 3774 3775 3800 3822 3846 3848

110479 220978 332019 442579 553402 610761 679382 818747 957312 1096364 1235494 1374912 1512799 1651995 1686701 1741747 1815928 1886823 1918083 2004536 2142256 2280976 2421367 2559885 2699105 2838849 2928741 3038489 3109590 3199925 3283757 3292933 3433300 3572423 3714467 3749106

Liters

40 80 120 160 200 *220.7 240 280 320 360 400 440 480 520 530 560 600 640 655.3 680 720 760 800 840 880 **920 960 1000 1040 1080 1117.1 1120 1160 1200 1240 ***1251.1

Weight (kg)

32.6 65.2 97.8 130.4 163.0 179.9 195.6 228.2 260.8 293.3 325.9 358.5 391.1 423.7 431.8 456.3 488.9 521.5 533.9 554.1 586.7 619.3 651.9 684.4 717.0 749.8 782.2 814.8 847.4 880.0 910.4 912.9 945.2 977.8 1010.4 1019.2

CG (mm) 3541 3547 3551 3552 3552 3553 3635 3754 3840 3910 3965 4012 4046 4079 4086 3993 3886 3785 3758 3785 3820 3853 3886 3913 3938 3961 3917 3901 3839 3804 3774 3775 3800 3822 3846 3848

Moment kg-mm

115437 231264 347288 463181 578976 639185 711006 856663 1001472 1146803 1292194 1438302 1582391 1728272 1764335 1822006 1899865 1973878 2079639 2097269 2241194 2386163 2533283 2678057 2823546 2969958 3063877 3178535 3253169 3347520 3435850 3446198 3591760 3737152 3885998 3921882

*Most critical fuel amount for most forward C.G. condition. **Most critical fuel amount for most aft C.G. condition at weight empties up to 3062 kilograms. ***Most critical fuel amount for most aft C.G. condition at weight empties at 3062 kilograms or greater. Weights given are nominal weights at 15°C.

(TABLE I.D. 910663)

Lateral (Engllh)

5-20


BHT-412-FM-4

Table 5-6.

Fuel loading table -

lateral (English)

USABLE FUEL LOADING TABLE (ENGLISH)


U.S. Gal.

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140.0 150 160 170 *173.1 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 330.5

Jet A, A-1, JP-5 (6.8 Lbs/Gal)

Weight (Lb)

CG In.

Moment In.-Lb.

U.S. Gal.

Weight (Lb)

CG In.

Moment In.-Lb.

65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1125 1170 1235 1300 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2148

0 0 0 0 0 0 -0.03 -0.06 -0.05 -0.04 -0.04 -0.03 -0.03 -0.03 -0.03 -0.43 -0.58 -0.69 -0.72 -0.69 -0.65 -0.62 -0.58 -0.56 -0.53 -0.51 -0.49 -0.48 -0.46 -0.44 -0.43 -0.41 -0.40 -0.39 -0.38 -0.38

0 0 0 0 0 0 -12 -27 -26 -23 -26 -21 -23 -25 -27 -419 -603 -762 -810 -807 -803 -806 -792 -801 -792 -796 -796 -811 -807 -801 -811 -800 -806 -811 -815 -816

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140.0 150 160 170 *173.1 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 330.5

68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1088 1156 1177 1224 1292 1360 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2108 2176 2244 2247

0 0 0 0 0 0 -0.03 -0.06 -0.05 -0.04 -0.04 -0.03 -0.03 -0.03 -0.03 -0.43 -0.58 -0.69 -0.72 -0.69 -0.65 -0.62 -0.58 -0.56 -0.53 -0.51 -0.49 -0.48 -0.46 -0.44 -0.43 -0.41 -0.40 -0.39 -0.38 -0.38

0 0 0 0 0 0 -12 -29 -27 -24 -27 -22 -24 -27 -29 -439 -631 -798 -847 -845 -840 -843 -828 -838 -829 -832 -833 -849 -845 -838 -848 -836 -843 -849 -853 -854

*Most critical fuel amount for left side most lateral C.G. condition.

(TABLE I.D. 910662)

5-21


BHT-412-FM-4

Table 5-6M.

Fuel loading table -

lateral (Metric)

USABLE FUEL LOADING TABLE (METRIC)

Jet A, A-1, JP-5 (.815 kg/I)

Jet B, JP-4 (.779 kg/I)

Liters

Weight (kg)

CG (mm)

Moment kg*mm

Liters

Weight (kg)

CG (mm)

40 80 120 160 200 220.7 240 280 320 360 400 440 480 520 530.0 560 600 640 *655.3 680 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 *1251.1

31.2 62.3 93.5 124.6 155.8 171.9 186.9 218.1 249.3 280.4 311.6 342.7 373.9 405.0 412.8 436.2 467.3 498.5 510.3 529.6 560.8 592.0 623.1 654.2 685.4 716.7 747.7 778.9 810.0 841.2 872.3 903.5 934.7 965.8 974.3

0 0 0 0 0 0 -1 -1 -1 -1 -1 -1 -1 -1 -1 -10 -14 -17 -18 -18 -17 -16 -15 -14 -13 -13 -12 -12 -11 -11 -11 -10 -10 -10 -10

0 0 0 0 0 0 -187 -218 -249 -220 -312 -343 -374 -405 -413 -4362 -6542 -8475 -9185 -9533 -9534 -9472 -9347 -9159 -8910 -9317 -8972 -8347 -8910 -9253 -9595 -9035 -9347 -9658 -9743

40 80 120 160 200 220.7 240 280 320 360 400 440 480 520 530.0 560 600 640 *655.3 680 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1251.1

32.6 65.2 97.8 130.4 163.0 179.9 195.6 228.2 260.8 293.3 325.9 358.5 391.1 423.7 432.0 456.3 488.9 521.5 533.9 554.1 586.7 619.3 651.9 684.4 717.0 749.8 782.2 814.8 847.4 880.0 913.0 945.2 977.8 1010.4 1019.2

0 0 0 0 0 0 -1 -1 -1 -1 -1 -1 -1 -1 -1 -10 -14 -17 -18 -18 -17 -16 -15 -14 -13 -13 -12 -12 -11 -11 -11 -10 -10 -10 -10

Moment kg*mm

0 0 0 0 0 0 -196 -228 -261 -293 -326 -359 -391 -424 -432 -4563 -6845 -8866 -9610 -9974 -9974 -9909 -9779 -9582 -9321 -9747 -9386 -9778 -9321 -9680 -10043 -9452 -9778 -10104 -10192

*Most critical fuel amount for left side most lateral C.G. condition. Weights given are nominal weights at 15°C.

NOTE This table Is invalid with auxiliary fuel tank(s) installed.

(TABLE

5-22

ID.

910671)

T N * Appendix A I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

BHT-412-FM-4

TABLE OF CONTENTS

Subject

Paragraph

Page Number

OPTIONALEQUIPMENT........................................

A-1 ..........

A-3

LIST OF TABLES

Title

Table Number

Flight Manual Supplements for Optional Equipment..........

A-1 ..........

Page Number A-3

A-1/A-2

BHT-412-FM-4

Appendix A OPTIONAL EQUIPMENT SUPPLEMENTS

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O A

Bell Helicopter Textron's policy is one of continuous product improvement and Bell reserves the right to incorporate design changes, make additions to and improve its products without imposing any obligation upon the company to furnish for or install such changes, additions, improvements, etc., on previously manufactured products.

The following items may be installed on the basic helicopter by authorized personnel. Only the optional equipment listed in this section require a Flight Manual supplement.

NOTE

Flight Manual Supplement numbers ending with .1 or -1 are applicable to Model 412 S/N 33001 — 33107 only. Supplement numbers ending with .2 or -2 are applicable to Model 412 S/N 33108 — 33213 and 36001 — 36019. Supplement numbers ending with .3 or -3 are applicable to Model 412 S/N 36020 — 36086. Flight Manual Supplement numbers ending with .4 or -4 are applicable to Model 412EP S/N 36087 and subsequent. Flight Manual Supplement numbers not ending with a decimal or dash number are applicable to all 412 helicopters including Model 412EP.

Table A-1. Flight Manual Supplements for Optional Equipment

NAME OF EQUIPMENT

KIT NUMBER

DATE CERTIFIED

CURRENT REVISION

BHT-412-FMS-1.4 Winterization Heater Operations

212-706-008

16 March 1995

Rev. 4 5 SEP 00

BHT-412-FMS-2 Emergency Floats

412-706-004

20 January 1981

Rev. 1 15 APR 96

BHT-412-FMS-3 Heated Windshield

412-706-010

20 January 1981

Reissued 8 OCT 94

BHT-412-FMS-4.1

Effectivity S/N 33001-33107

BHT-412-FMS-5.1


BHT-412-FMS-6 Flight Director

412-706-111

13 February 1981

Reissued 8 MAY 89

23 MAY 2002

Rev. 23

A-3

BHT-412-FM-4 Table A-1. Flight Manual Supplements for Optional Equipment (Cont) NAME OF EQUIPMENT

KIT NUMBER

DATE CERTIFIED

CURRENT REVISION

BHT-412-FMS-7 Internal Hoist

214-706-003

2 October 1981

Reissued 11 MAY 95

BHT-412-FMS-8 Litter Kit Operations

412-706-006

29 September 1981

Reissued 5 OCT 94

BHT-412-FMS-9.4 External Cargo Operation

212-706-103

25 September 1991

Reissued 15 SEP 95

BHT-412-FMS-10.3 Category A Operations for Helicopters Equipped with PT6T-BE Engines

Effectivity S/N 36020 — 36086

BHT-412-FMS-CAA-10.4 Group A Operations

412-705-018

11 November 1996

Original

BHT-412-FMS-11.1 20/20 Auxiliary Fuel


BHT-412-FMS-12 Nightsun Searchlight

212-899-333/ 412-899-267

04 December 1981

Reissued 08 MAY 89

BHT-412-FMS-13 Cold Weather Operations

412-703-004

Data incorporated into Basic Flight Manual

BHT-412-FMS-14 Thailand Special Avionics

412-899-003

13 February 1982

Not Printed

BHT-412-FMS-15 Fixed Step

212-706-057

06 February 1982

Reissued 23 JUN 94

BHT-412-FMS-16 Droop Restraint, Main Rotor and Stick Centering Indicator

412-704-114/ 412-704-115

Data incorporated into Basic Flight Manual

BHT-412-FMS-17.4 Auxiliary Fuel Supplement

412-706-007

05 January 1984


A-4

Rev. 23

23 MAY 2002

Reissued 23 JUN 94


BHT-412-FM-4

Table A-1. Flight Manual Supplements for Optional Equipment (Cont) DATE CERTIFIED

CURRENT REVISION

NAME OF EQUIPMENT

KIT NUMBER

BHT-412-FMS-18.3 Loadhailer Operations


BHT-412-FMS-19.4 Soft Interior

412-705-510

28 March 1985

Reissued 5 OCT 94

BHT-412-FMS-20 Weather Radar Kit

412-899-107

16 June 1986

Reissued 5 OCT 94

BHT-412-FMS-21 Global Navigation System, GNS-500A/S3 with NAV Switching

412-899-141

16 June 1986

Reissued 8 MAY 89

BHT-412-FMS-22.2 Category A Operations for Helicopters Equipped with PT6T-3BEngines

Effectivity S/N 33108- 33213 And S/N 36001 - 36019

BHT-412-FMS-22.3 Category A Operations for Helicopters Equipped with PT6T-3BEEngines

Effectivity S/N 3 36020 -

BHT-412-FMS-23

Reserved

BHT-412-FMS-24 Seat Cushion Kit

412-706-019

24 July 1987

Reissued 8 DEC 95

36086

412-706-009 BHT-412-FMS-25.4 Auxiliary Fuel Operations

10 March 1988

Reissued 23 JUN 94

BHT-412-FMS-26 Two-Speed Internal Hoist

412-899-223/ 214-706-003

19 September 1988

Reissued 11 MAY 95

BHT-412-FMS-27 Litter Kit Operation

205-706-047

14 October 1988

Reissued 23 JUN 94

BHT-412-FMS-28.3 Dual Battery


BHT-412-FMS-29.3 Removal of Upper Aft Center Fuel Cell


9 November 1989

Reissue 23 JUN 94

BHT-412-FMS-CAN-30 Canadian Addendum to the Supplements for Internal Hoist and External Cargo Operation

Rev. 18

A-5


KIT NUMBER

BHT-412-FMS-31 Category B Operations with Approved Configuration of Nine or Less Passenger Seats

STC No. SH7727SW

DATE CERTIFIED

CURRENT REVISION

08 February 1990

Original


BHT-412-FMS-32.2

Not Printed

BHT-412-FMS-33 Loran C Navigation System (King KLN-88)

412-899-231

BHT-412-FMS-34.2 Improved Hover Performance with PT6T3BE Engines and 5-Minute Takeoff Power Rating

Effectivity S/N 33108-33213 and 36001-36019

BHT-412-FMS-35.4 Category B Operations When Configured with Nine or Less Passenger Seats

22 June 1990

Original

10 April 1991

Rev. 3 5 SEP 00

BHT-412-FMS-36.3

Not Printed

BHT-412-FMS-37.4 Dual Flight Director (3-Axis)

412-706-024

18 October 1994

Rev. 1 20 DEC 99

BHT-412-FMS-38.4 EFIS with Four-Axis AFCS Kit 412-705-009 AND 412-705-024 with or without Increased MC Power Kit

412-706-029

18 October 1994

Rev. 2 16 AUG 00

BHT-412-FMS-39.4 412-570-002 Dual Digital Automatic Flight Control System, Search and Rescue (SAR) with PT6T-3B/E or PT6T-3D Engines

11 September 1992

Rev. 3 26 MAR 98

BHT-412-FMS-40 Increased Generator Capacity Kit

29 October 1992

Reissued 05 OCT 94

A-6

Rev. 23

23 MAY 2002

412-706-026

BHT-412-FM-4 Table A-1. Flight Manual Supplements for Optional Equipment (Cont) DATE CERTIFIED

CURRENT REVISION

NAME OF EQUIPMENT

KIT NUMBER

BHT-412-FMS-41.3 Increased Maximum Continuous Power Kit with PT6T-3BE Engines


BHT-412-FMS-42

Reserved

BHT-412-FMS-43.4 Special Search and Rescue Configuration

412-899-052

08 March 1993

Reissued 02 FEB 95

BHT-412-FMS-44.4 Airline Passenger Seating

412-706-028

18 June 1993

Rev. 1 13 JUL 99

BHT-412-FMS-45.4 Increased Maximum Continuous Power Kit with PT6T-3D Engine

412-706-029

27 August 1993

Rev. 3 5 SEP 00

BHT-412-FMS-46.3 PT6T-3D Engines


BHT-412-FMS-47 Folding Step

412-899-287

25 October 1993

Original

BHT-412-FMS-48.1 and 48.2 Engine No. 2 Gov Trim Switch

Effectivity S/N 33001-33213 S/N 36001-36019

BHT-412-FMS-49.4 Honeywell Primus 700 Weather Radar

412-705-034

05 October 1994

Original

BHT-412-FMS-50 through 52

Reserved

BHT-412-FMS-53.3 and 53.4 PT6T-3DE (30 Minute OEI Rating)

412-706-050

29 March 1996

Reissued 21 JAN 99

BHT-412-FMS-54

Reserved

BHT-412-FMS-55.4 Emergency Floats with Automatic Retracting Passenger Step

412-706-004 Effectivity S/N 36122, 36123 S/N 36125 and subsequent

15 April 1996

Reissued 31 MAY 97


23 MAY 2002

Rev. 23

A-7


KIT NUMBER

DATE CERTIFIED

CURRENT REVISION

BHT-412-FMS-56.3 and 56.4 PT6T-3DF Engine (30 Minute OEI Rating)

412-706-051

08 August 1996

Original

BHT-412-FMS-CAA-57.4 United Kingdom Registered Helicopters

412-705-016/ 412-705-018

11 November 1996

Rev. 1 19 MAR 98

12 March 1997

Original


BHT-412-FMS-58.4 External Class D Loads

BHT-412-FMS-59.4 Trimble 2101 GPS With Electrical/Mechanical Display For VFR Operations

412-899-510

17 June 1997

Original

BHT-412-FMS-60.4 Customizing Royal Thai Air Force

412-899-095

23 December 1997

Reissued 10 FEB 98

BHT-412-FMS-61.4 Slope Landing Operations

412-704-012

09 July 1997

Reissued 29 JUL 97

23 April 1999

Rev. 1 02 JUN 99

19 September 1997

Rev. 1 22 OCT 97

BHT-412-FMS-62.4 Category A Operations PT6T-3D Engine BHT-412-FMS-63.4 Self Sealing Fuel Cells

412-899-175

BHT-412-FMS-64.4

Reserved

BHT-412-FMS-65.4 Ten Cell — Self Sealing Fuel

412-899-377

22 June 1998

Rev. 1 02 JUL 98

BHT-412-FMS-66.4 Global Positioning System (Trimble 2101 Approach) with Electrical/ Mechanical HSI Display

412-899-378

19 June 1998

Original

A-8

Rev. 23

23 MAY 2002


KIT NUMBER

BHT-412-FMS-67.1 and 67.2 PT6T-3BF Engines (30 Minute OEI Rating)

Effectivity S/N 33001 and Subsequent S/N 36001 — 36019

BHT-412-FMS-68.3 PT6T-3BG Engines (30 Minute OEI Rating)

Effectivity S/N 36020 — 36086

DATE CERTIFIED

CURRENT REVISION

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O 412-705-008 BHT-412-FMS-69.4 Dual Digital Automatic Flight Control System, Search and Rescue

29 January 1999

Original

BHT-412-FMS-70.4 Flight Management System, UNS - 1C/D

412-899-381

29 January 1999

Reissue 27 APR 00

BHT-412-FMS-71.4 External Hoist (Breeze-Eastern)

412-709-001

18 August 2000

Original

BHT-412-FMS-72.4 Cockpit Voice Recorder

412-705-001-115 23 May 2002 or 412-704-026-101

23 MAY 2002

Original

Rev. 23

A-9/A-10

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-1.4

ROTORCRAFT FLIGHT MANUAL SUPPLEMENT

WINTERIZATION HEATER OPERATIONS (212-706-008) 36087 AND SUB CERTIFIED 16 MARCH 1995

This supplement shall be attached to the Model 412 EP Flight Manual (BHT-412-FM-4) when the 212-706-008 Winterization Heater has been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, or other applicable supplements, consult basic Flight Manual.

__________ NOTICE COPYRIGHT COPYRIGHT 2000 ® INC. TEXTRON BELL HELICOPTER TEXTRON INC ANDBELLHELICOPTER

A DIVISION OFTEXTRON CANADA LTD. ALL RIGHTS RESERVED

Bell Helicopter

Inj:.*It

of Textron Inc. A Subsidiary 76101 BOX482 - FORTWORTHTEXAS POSTOFFICE . i

16

MARCH 1995

REVISION 4 - 5 SEPTEMBER 2000


FAA APPROVED

NOTICE PAGE

Additional copies of this publication may be obtained by contacting: Commercial Publication Distribution Center Bell Helicopter Textron Inc. P. 0. Box 482 Fort Worth, Texas 76101-0482

NP

BHT-412-FMS-1.4

T N 1ell î I R P N O R (P9HT I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODEL

SW/Il/A

ROTOR CRAFT

MANUAL

W

COPLEMENT a EZATION HEATER TIONS 6-008) <<27

36087 ADNT

TEMPORAN MAST TORQUE SIGN LTIONER / ASB 412-91

nt pubis This Temporary Revision extends performance and provides new performnbe data compliance with ASB 412-97-90. Attach these temporary revision pages next to lik pages in Flight Manual after compliance with ASB 41\ DO NOT remove existing pages from Flight Manual.

Information contained herein supplements information ofb Fight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Right Manual.


Bell HelicopterEi :ElEi ] A SuTifary of TEltiIn

COPYRIQIT NOTICE

COIIUIT_

20

I00 C

CPTO TTmoT INC. BOil A BEL HELICOPER [ XTRON, A DIVISION OFTEIR CANADLTD IWALL fIsREED m

1xS

Io0I E1

2

.W WOtHn FORT

L 7&101

16 MARCH 1995


3

1


LOG OF REVISIONS

Original........... 0 ............. 16 MAR 95 Revision..........1 ............. 29 AUG 95 Revision .......... 2 .............. 13 JAN 97

Revision ......... 3 ............. 08 AUG 97 Revision .......... 4 ............. 05 SEP 00

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title ..................................... NP ......................................

4 0

REVISION NO.

PAGE

A - B .................................. 4 0 . ..................................... i/ii . 1- 2...................................0 3-18 .................................. 4

a_^^~~~ ~NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

SEP O)2000'

Rev. 4

A



Temporary ......... ..... 8 AUG 97 Canceled ....................... 5 SEP 00

NOTICE

This Temporary Revision has been cancelled. Subject data has been incorporated in BHT-412-FMS-1.4 at Revision 4. Remove and discard all Temporary pages dated 8 August 1997 from this manual including these canceled pages.

APPROVED:

;

M ANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170 TEMPORARY REVISION - 8 AUGUST 1997 B CANCELED - 5 SEP 2000

'SEP

5 2000


LOG OF FAA APPROVED REVISIONS

Original ........... 0 ............. 16 MAR 95 Revision .......... ............. 29 AUG 95 Revision .......... 2 .............. 13 JAN 97

DATE:

APPROVED:

ROTORCRT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

B

Rev. 4

Revision .......... 3 ............. 08 AUG 97 Revision .......... 4 ............. 05 SEP 00

SEP 05 2000

EP

5 2000

T


FAA APPROVED

GENERAL INFORMATION

The Winterization Heater provides greater cabin heating capacity than the basic heater. The Winterization Heater comprises a larger mixing valve, a larger noise

suppressor, extra outlets, ducts, and the hardware necessary for installation. Operation of the Winterization Heater is identical to the basic heater.

T N I Section 1 R P N O R I O T F C U T O OD N R E Section 2 P E E R R C S N G O N I Section 3 R W O E I F V NLY O FAA APPROVED

1-6.

BHT-412-FMS-1.4

WEIGHT AND CENTER OF *„--------GRAVITY -weight

readjusted if necessary to return empty CG within allowable limits.

Actual weight change shall be determined after the heater is installed and ballast

No change from basic manual.


T N I Section 4 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-1.4

4-1.

INTRODUCTION

Performance with Winterization Heater switched OFF is same as that shown in basic Flight Manual for heater off. Performance with Winterization Heater switched ON is presented as follows:

Refer to figure 4-1 for out-of-groundeffect hover performance.

2

FAA APPROVED

In-ground-effect hover performance is same as that shown in basic Flight Manual for Heater ON. Refer to figure 4-2 for takeoff Refer to e e performance. R t f Refr to fi performance.

4 e 43

fr

c


BHT-412-FMS-1.4

HOVER CEILING


SKID HEIGHT 60 FT. TAKEOFF POWER WINTERIZATION HEATER ON ENGINE RPM 100% GENERATOR 150 AMPS(EA.) -40 TO 20°C CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

14,000 FT Ho LIMIT

LoZ/I _ _I

I

7 /Ik , J/__/,&^gP /

/_

-40

-20

0

OAT - "C

I_

I,

20

AT I'

I

°C

__843 ^" _4___/_ 0 2o,

_

/,

I

8

____

9

10

GROSS WEIGHT -

11

12

LBS x 1000

0

5.0 5.4 4.2 4.6 3.4 3.8 1000 4 Fgr Hovr /_GROSS WEIGHT - KG X X 1000 _________GROSS ^&b~~~~ ^^^b~ ~ Figure 4-1. Hover ceiling - out of ground effect (Sheet 1 of 2)

!SEP n5

n.

Rev./

T N I R P N O R I O T F C U /'YSX~~~~~P ! T /! O OD N R E P E E R R C S …CN G O N I R W O E I F V NLY O BHT-412-FMS-1.4

FAA APPROVED

HOVER CEILING





14,000 FT H0 UMIT

-40

AT -

-40

-20

0

12

1-1

1000 WEIGHT -- KGS X 1000 GROSS WEIGHT GRSS

C

OAT - C

1-0

9

8

20

0

-20

20

8

9

10

GROSS WEIGHT I

I

I

I

I

I

11

12

LBS x 1000

I

I

I

I

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

Figure 4-1.

4

Rev. 4

Hover ceiling - out of ground effect (Sheet 2)

SEP 0520 0 0

0


BHT-412-FMS-1.4

TAKEOFF DISTANCE OVER 50 FT. OBSTACLE

INITIATED FROM 4 FT SKID HEIGHT VTOCS = 45 KIAS WINTERIZATION HEATER ON

HOVER POWER + 15% TORQUE ENGINE RPM 100% GENERATOR 150 AMPS EACH

14.000 FT. DEN. ALT. LIMIT

w--

1

^\

lIi/-~1------I-I-

MI--

L

I

MIN-l OAT

1a/ / --1YII- -1~AIIGROSS -'

FOR TAKEOFF

t

_!

oO

0

_

-40-20

OAT-

0

_MAXIMUM WEIGHT

20

OC

200

400

600

800

1000

1200

TAKEOFF DISTANCE - FEET

ENGLISH _U Z-~ENGLISH ^412FMS1.4-4-2-1

Figure 4-2. Takeoff distance (Sheet 1 of 2)

SEP 05 2000

Rev. 4

5


FAA APPROVED

TAKEOFF DISTANCE OVER 15 METER OBSTACLE

HOVER POWER + 15% TORQUE ENGINE RPM 100% GENERATOR 150 AMPS EACH

INITIATED FROM 1.2m SKID HEIGHT VTOCS = 45 KIAS WINTERIZATION HEATER ON

j

4 2 67

METERS DEN. ALT. LIMIT

\IS/& I- !

°

I__\\

I

.{

LMAXIMUM / /_ ____ GROSS WEIGHT

TA!4

Ml

1_JHLZLLFOR TAKFF

1

_

T___l

^2

dI //

X UTVI

|11OAT-h

-40-20 0 20 OAT - °C

J/1I'

ll

rV___[ o/

/

-1

CO

100 150 200 250 300 TAKEOFF DISTANCE - METERS METRIC

Figure 4-2.

6

Rev.4

SEP 052000

-:

Takeoff distance (Sheet 2)

412FMS1.4-4-2-2


BHT-412-FMS-1.4

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 7000 LB (3175 kg)

70 KIAS WINTERIZATION HEATER ON





0

20000

I

i ii

-hi

i

18000

i

i

6

-

-~______\~ __ _

16000

_______ -_

\

5

_

140 t 14000

__________

0

OFCL-4 RATE-_

1,ooo ^^^^-^\^^ I.

0

0

c

~a L

igrJ 4-3

0

;So

Twin enierit-e

8000

)RATE

60004-3. 4000 2000

0

-2 Figure

400

OF CLIMB

FT/MIN

-

n

Twin engine r ate of climb (Sheet S 1 of 1 2)

800

Figure 4-3.

1600 2000 2400 2800 RATE OF CLIMB - FT/MIN 1200

3200 3600

Twin engine rate of climb (Sheet 1 of 12)

SEP0 5 200

,Rev. 4

7


FAA APPROVED

TWIN ENGINE RATE OF CUMB GROSS WEIGHT 8000 LB (3629 kg)





RATE OF CLIMB - M/SEC

0 1 2

20000

.

3 4

5

6 7

iJ

.I

8 9 10 11 12 13 14 15 16 17 18 ll ll l mll il

'

18000

16000

5

5 14000 ,

L_

o

-RATE OF CLIMB

Q

2000 800000

-

0

--------

400

-

800

Figure 4-3.

8

Rev. 4

SEP 05 2000

-

-

-002

1200 1600 2000 RATE OF CLIMB -

FT/

0

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2400 2800 FT/MIN


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0



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1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB - FT/MIN Twin engine rate of climb (Sheet 3)

.SEP 05

O

Rev. 4

9

T N I R P N O R I O T F C U 0 T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-1.4

FAA APPROVED

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB (4536 kg)



1. CUMB SPEED IS 60 KIAS 2. RATE OF CUMB WILL DECREASE 275 FTIMIN


RATE OF CLIMB - M/SEC 0 1 2 3 4 5 6 7 8 9 101112 131415161718

20000

Uli6il

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18000

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SEP 052000


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SEP 052000

Rev. 4

11


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1 2 3

0

20000

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RATE OF CLIMB - M/SEC 4 5 6 7 8 9 10 11 12 13 14 15 1617 18 mili

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1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

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2 Rv.

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BHT-412-FMS-1.4







20000

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13


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RATE OF CLIMB -

0

20000

M/SEC

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FT/MIN

RATE OF CLIMB -

FT/MIN

Figure 4-3. engine rate of climb (Sheet 8) 4 4 0 000 2000Twin 400 5 800 1200 1600 2000 2400 2800 Figure 4-3.

14

Rev. 4

--

_

SEP 052000


3200

3600 4

T N I R P N O 0^yb~R I O T F C U T O OD N R E P E E R R C S N G O / N I R W O E I F Y V L 0 N O FAA APPROVED

BHT-412-FMS-1.4



70 KIAS


WINTERIZATION HEATER ON



~~RATE

0

1 2

N

20000

" Hil

OF CLIMB - M/SEC

3 4 I

5

6

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9 10 11 12 13 14 15 16 17 18 ill m ml l ul i I ii

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1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

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FT/MIN

Twin enginie rate of climb (Sheet 9)

SEP 052000

Rev.4

15


FAA APPROVED





0

20000 20000

1 2


RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 1011 12 131415161718 l ihlll

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Rev. 4

17


FAA APPROVED






RATE OF CLIMB -

0

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4 5 6

7 8

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9 10 11 12 13 14 15 16 17 18

20000

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FT/MIN


T N I BELL R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-2

MODELS


EMERGENCY FLOATS (412-706-004) CERTIFIED 20 JANUARY 1981

This supplement shall be attached to the

Models 412 and 412 EP Flight Manual when the 412-706-004 Emergency Floats have been installed. Refer to BHT-412-FMS-55.4

for S/N 36122,

36123, 36125 and subsequent.

Information contained herein supplements information of basic Flight Manual. For

Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter TEXTRON NOTICE COPYRIGHT COPYRIGHT 1996 INC. BELL HELICOPTER ANDBELLHELICOPTER TEXTRON INC A DIVISION OFTEXTRON CANADA LTD.

ALLRIGHTS RESERVED

of Textron Inc A Subsidiary

TEXAS 76101 BOX482* FORTWORTH, POSTOFFICE

23 JUNE 1994 23

REVISION 1 -15

JUNE

1994

APRIL 1996

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-2

NOTICE PAGE

Additional copies of this publication may be obtained by contacting: Commercial Publication Distribution Center Bell Helicopter Textron P. 0. Box 482 Fort Worth, Texas 76101-0482

NP



FAA APPROVED

LOGOFAPPROVED REVISIONS

Original ........... Reissue...........0 Reissue...........0

0 ............... .............. ...............

20 Jan 81 16 May 89 20 Oct 93

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT. WORTH,TX 76193-0170

B

Reissue ........... Revision ..........

0 .............. 1 ...............

23 Jun 94 15 Apr 96


BHT-412-FMS-2

GENERAL INFORMATION

The emergency floats will allow the helicopter to land in water during an emergency situation. The kit consists of six skid mounted float bags, four water activated float switches that are mounted on the underside of the fuselage, a pneumatic bottle located in the nose compartment, a collective mounted EMERG FLTS switch, a pedestal mounted EMER INFLATIONhandle, and the necessary hardware to complete the installation. Also, the kit includes provisions for installing life rafts and vests. The floats automatically inflate when the EMERGFLTSswitch Is in the ARMED position and the water activated float switches are submerged. If the system or one of the float bags rupture after inflation, check valves in each float bag will maintain float integrity. An instrument panel mounted EMER FLOATS caution light illuminates whenever the EMERG FLTS switch is in the ARMED position. System security is protected by an EMERG FLOATScircuit breaker located on the overhead console. The EMERINFLATIONhandle provides for manual activation of the emergency flotation system in the event the floats fail to inflate automatically.

i/ii


BHT-412-FMS-2

Section

1

TYPE OF OPERATION The emergency

floats

FORWARD

are installed

for

assistance during emergency ditching and use is approved for VFR and IFR operation. FLOATS switch shall be OFF and FLOATS

FLIGHT

OR DESCENT.

REFER TO INADVERTENT FLOAT INFLATION IN FLIGHT (SECTION

3).

Maximum autorotation stowed):

airspeed (floats

caution light extinguished when operating over land.

105 knots below 10,000 feet pressure altitude

Emergency float safety pin shall be removed prior to flight over water.

80 knots above 10,000 feet pressure altitude

FLIGHT WITH OPTIONAL EQUIPMENT INSTALLED

Maximum forward speed for ditching: 33 knots in calm water

Fixed passenger step shall not be installed concurrently with emergency floats.

Retractable passenger step shall be stowed prior to flight over water.

When internal hoist is installed, hoist cable anti-chafing guard shall be installed on same side of helicopter as hoist.

15 knots in rough water

Rate of descent should be reduced as low

as possible upon water contact

ALTITUDE LIMITATIONS

ALTITUDELIMITATIONS

Maximum pressure altitude for Inflation of emergency floats is 10,000 feet. Helicopter

operation above 10,000 feet is permitted provided the FLOATSswitch is in the OFF

AIRSPEED AIRSPEED LIMITATIONS LIMITATIONS

Inflation of emergency floats during

forward flight is prohibited. Flight after landing with floats Inflated is prohibited.

position and FLOATScaution light is extinguished. extinguished.

WEIGHT/CGLIMITATIONS

Actual weight change shall be determined

after floats are installed and ballast readjusted,

SEVERE NOSE UP PITCHING WILL OCCUR IF EMERGENCY

FLOATS ARE INFLATED

IN

If necessary, to return empty

weight cg within allowable limits.

The emergency float kit does not change

the CG limits of the helicopter.

1

T N I R P N O R I O T Section 2 F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-2

FAA APPROVED

FLOAT ARMING Floats shall

armed: 1. 2. 3.

not be

PLACARDS

FLOAT INFLATION OR OPERATION IN FORWARD FLIGHT IS PROHIBITED

Over land. Above 60 KIAS. Above 600 feet AGL.

(Located above pilot airspeed indicator.)

EXTERIORCHECK

BEFORE TAKEOFF

OVER

WATER

NOTE

STEP switch -

Ensure that emergency floats have had periodic

inflation

FLOATS switch -

and

ARMED; check FLOATS

caution light Illuminated.

inspection.

Emergency floats -

STOW.

IN-FLIGHT OPERATION

Stowed.

Emergency float covers and supports -

feet above water and airspeed of 60 KIAS

or above, or when over land; check

INTERIOR CHECK

FLOATS caution light extinguished.

EMER INFLATION handle

-

safetied.

Down and

Nitrogen bottle - Secured and pressure within allowable limits for pressure altitude

and

ambient

temperature

as

shown on chart decal, located on nitrogen bottle. Emergency floats safety pin over-water flight.

Remove for

BEFORE

WATER

LANDING

OVER

FLOATSswitch - ARMEDwhen below 600 feet above water and airspeed of 60 KIAS

or below; check FLOATS caution light illuminated.

AFTER LANDING

PRESTARTCHECK

FLOATS switch - OFF; check FLOATS caution light extinguished.

FLOATS switch - OFF.

BEFORE

EM'R FLOATScircuit breaker - IN.

HELICOPTER

LEAVING

Emergencyfloats safety pin 2

Installed.

T N I Section 3 PR N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-2

INADVERTENT

FLOAT

INFLATIONIN FLIGHT

In event of unintended inflation, reduce airspeed, adjust rate of descent to 200 fpm or less, and land as soon as possible. Avoid landing on terrain which could damage float bags.

ELECTRICALMALFUNCTIONS

FLOATS CAUTIONLIGHT ON WITH FLOATSSWITCHOFF

EMERGENCY DITCHING WARNING

CREW AND PASSENGERDOORS SHALL REMAINCLOSED DURING DITCHING.

FLOATSswitch - ARMED;check FLOATS caution light illuminated. Establish an autorotative glide speed of 65 to 70 knots for all gross weight.

If FLOATS caution light illuminates with FLOATS switch OFF, pull EMER FLOATS circuit breaker. NOTE

If EMERFLOATScircuit breaker is pulled floats will not deploy automatically.

If FLOATS caution light remains illuminated after pulling circuit breaker, land as soon as practical.

ELECTRICALSYSTEMFAILURE

Pull EMER INFLATION handle upon water contact to inflate floats.

SEVERE NOSE UP PITCHING WILL OCCUR IF EMERGENCY FLOATS ARE INFLATED IN FORWARDFLIGHTOR DESCENT.

At 100 feet altitude execute a moderate cyclic flare to reduce airspeed. Adjust collective and cyclic pitch sufficiently to touchdown in a nose-up attitude with forward speed and rate of descent as low as possible. If emergency floats do not inflate immediately upon water contact, pull EMERINFLATIONhandle.

After water touchdown, complete shutdown, check for damage, and determine if helicopter should be abandoned. 3

T N I R P N O R I O T F C Section4 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-2

FAA APPROVED

NOTE

Crew egress should be accomplished by jettisoning crew doors.

Takeoff after emergency landing with floats inflated is prohibited.

EGRESSAFTER DITCHING

Passenger egress should be accomplished by pushing out passenger door windows at corners marked EMERGENCYEXIT PUSH HERE.

Ensure rotor is stopped prior to egress.


T N I R P N O R I O T F C ROTORCRAFT U T D MANUAL N O OFLIGHT R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-3

MODELS

SUPPLEMENTFOR HEATEDWINDSHIELD (412-706-010)

CERTIFIED 20 JANUARY 1981

This supplement shall be attached to the Models 412 and 412 EP Flight Manual when the 412-706-010 Heated Windshield has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter NOTICE COPYRIGHT COPYRIGHT 1994 INC TEXTRON BELLHELICOPTER TEXTRON. HELICOPTER ANDBELL LTD OFTEXTRON CANADA ADIVISION ALLRIGHTS RESERVED

A Subsidiary of TextronInc POST OFFICE BOX 482 * FORT WORTH. TEXAS 76101

REISSUE- 18 OCTOBER1994

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL

Log of Pages

FAA APPROVED SUPPLEMENT

BHT-412-FMS-3

LOG OF REVISIONS

Original ..... Reissue .... Reissue .....

0 0 0

............. ............. .............

20 Jan 81 08 May 89 18 Oct 94

LOG OF PAGES

Revision No.

Page

Title - A ................... i/ii ........................ 1-2 ..................... 3/4 .......................

Page

0 0 0 0

APPROVED:

MANAGER


NOTE: Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

A

Revision No.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

412 ROTORCRAFT FLIGHT MANUAL

INTRODUCTION BHT-412-FMS-3

INTRODUCTION

The Heated Windshield Kit allows the crew to electrically defrost/defog the windshield. The kit consists of two heated windshield panels, switches, circuit breakers. caution/advisory lights, and the necessary hardware and wiring to complete the installation.

The ON/OFF switches are located on the overhead console and identified as WSHLD HEAT LH and RH. Placing either switch in the ON position activates the corresponding heated windshield panel and illuminates the respective ON advisory light. The caution/advisory lights are located on the instrument panel. The caution portion of the light illuminates HOT whenever the respective heated windshield panel overheats. Electrical power is provided by the 28 Vdc NON ESNTL bus. CONT and PWR circuit breakers, located on the overhead console and identified as WINDSHIELD HEAT LH and RH. provide circuit protection.

i/ii

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

412 ROTORCRAFT FLIGHTMANUAL

Section1

BHT-412-FMS-3

PLACARDSAND DECALS

CAUTION

STAND-BY COMPASS UNRELIABLE WITH WINDSHIELD HEAT ON

WSHLD HEAT

LH

RH

INDICATOR

ON

HOT

ON

HOT

1

T N I Section 2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL

Section 2


BHT-412-FMS-3

PRESTART CHECK

WSHLD HEAT LH switch - OFF.

BEFORE TAKEOFF

WSHLD HEAT LH switch - ON (if desired); check ON light illuminates.

WSHLD HEAT RH switch - OFF.

WINDSHIELD HEAT circuit breakers - In. BATTERY BUS 1 switch - ON.

WINDSHIELD HEAT caution/advisory light segments - PRESSTO TEST: check ON lights illuminate green, HOT lights illuminate amber. NOTE

The intensity of the heated windshield lights is controlled by the caution panel BRIGHT/DIM switch when PILOT INSTR LT rheostat is on.

2

WSHLD HEAT RH switch - ON (if desired): check ON light illuminates. NOTE

The windshield heat green advisory lights will illuminate and extinguish as the windshield heat cycles on and off during normal operation.



Sections3 &4

T N I * Section 3 R P N O R I O T F C U T O OD N R E P E E R R C S Section 4 N G O N I R W O E I F V NLY O BHT-412-FMS-3

DC GENERATORFAILURE

NON-ESNTL BUS switch - MANUAL. NOTE

DC power for the heated windshield is supplied by the nonessential bus.

WINDSHIELD BREAKER

HEAT

CIRCUIT

If LH or RH WINDSHIELD HEAT CONT or PWR circuit breaker pops out, turn respective WSHLD HEAT switch OFF.

WINDSHIELD HEAT CAUTION LIGHT

If either WSHLD HEAT caution light illuminates (HOT), turn respective WSHLD HEAT switch OFF.

No change from basic Flight Manual.

3/4

T N I R P N 3ell Model 412 O R I O T F C ROTORCRAFT U T D MANUAL N O OFLIGHT R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-4.1

33001 THROUGH33107

SUPPLEMENTFOR AUXILIARY FUELOPERATIONS (412-706-007)

CERTIFIED FEBRUARY10, 1981

This supplement shall be attached to the Model 412 Flight Manual (S/N 33001 through 33107) when the 412-706-007 Auxiliary Fuel Kit has beeninstalled

The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter A Subsidiary of Textion Inc POST OFFICE BOX 482 · FORT WORTH. TEXAS 76101

10 FEBRUARY1981

REVISION3 - 8 FEBRUARY1985

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W 0 O E I F V NLY O FAA APPROVED SUPPLEMENT


BHT-412-FMS-4.1

LOGOF REVISIONS

Original. Revision .... Revision .... Revision ....

.... .10 Feb81 1 .... 09Sep81 2 .... 08 Sep 83 3 .... 08 Feb85

LOGOF PAGES

FLIGHT MANUAL

MANUFACTURER'S

Revision No.

Page

Title....................... A......................... i/ii ........................ 1-1/1-2.................... 2-1/2-2.................... 3-1/3-2.................... 4-1/4-2 ....................

3 3 3 3 3 3 3

Page

1-1 - 1-12................. 1-13/1-14..................

APPROVED:

MANAGER

AIRCRAFTCERTIFICATIONDIVISION FEDERALAVIATION ADMINISTRATION DEPARTMENTOF TRANSPORTATION SOUTHWESTREGION,FORTWORTH, TEXAS

NOTE: Revisedtext is indicated by a black vertical line. Insert latest revision pages; dispose of supersededpages.

A

Rev. 3

DATA

Revision No. 3 3



INTRODUCTION

BHT-412-FMS-4.1

INTRODUCTION

The Auxiliary Fuel Kit provides additional fuel capacity to extend the range of the helicopter. The kit consists of a left and right auxiliary fuel tank and the hardware and wiring necessary to complete the installation. The left or right auxiliary fuel may be removed as operational requirements dictate.

One fuel tank provides an additional 81.5 U.S. gallons (308.5 liters) of fuel. Both fuel tanks combined, provide an additional 163 U.S. gallons (617 liters) of fuel.

Rev. 3

i/ii



Section 1

BHT-412-FMS-4.1 I

WEIGHT/CG LIMITATIONS

Actual weight changes shall be determined after installation of auxiliary fuel tank(s), and ballast shall be readjusted, if necessary, to return empty weight CG within allowable limits.

The gross weight center of gravity limits as presented in the basic Flight Manual do not change when either or both auxiliary fuel tanks are installed. Refer to Manufacturer's Data. Section 1, for weight and balance data.

Rev.3

1-1/1-2

T N I Section 2 PR N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT


Section 2

BHT-412-FMS-4.1 I

ge from basic Flight Manual.

0

Rev. 3

2-1/2-2


Section 3


BHT-412-FMS-4.1 I


Rev. 3

3-1/3-2

T N I Section 4 PR N O R I eNo FO cha T C U T O OD N R E P E E R R C S N G O N I R W 0 O E I F Y V L 0 N O FAA APPROVED SUPPLEMENT

Section 4


BHT-412-FMS-4.1 I

ige from basic Flight Manual.

0

Rev. 3

4-1/4-2

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SUPPLEMENT

412 ROTORCRAFT MANUFACTURER'S DATA

Section 1

BHT-412-FMS-4.1

MANUFACTURER'SDATA

AUXILIARY

FUEL

SYSTEM

AUXILIARY FUEL SYSTEM SERVICING

The auxiliary fuel tanks are interconnected with the basic fuel system to allow gravity flow of auxiliary fuel into main fuel cells as fuel is consumed. The auxiliary fuel system is serviced simultaneously with the basic fuel system through the single filler port located on the aft right side of the fuselage. FUEL SYSTEM CAPACITIES

BASIC SYSTEM WITH AUXILIARY TANK

LEFT OR RIGHT

Total capacity: 295.8 U.S. gallons (1119.7 liters) Usablefuel: 293.0 U.S. gallons (1109.1 liters)

BASIC SYSTEM WITH TANKS

BOTH AUXILIARY

Total capacity: 377.5 U.S. gallons (1429.0 liters) Usable fuel: 374.7 U.S. gallons (1418.4 liters)

AUXILIARY FUEL LOADING TABLES

Fuel loading tables are presented for weight and balance computations in both English and Metric units. These tables shall be used in lieu of the tables for the basic fuel system when either or both auxiliary fuel tanks are installed. Weightsand moments listed herein represent usable fuel on board to include that contained in basic fuel cells. Refer to table 1-1 (English)or 1-1M (Metric) when both auxiliary tanks are installed. Tables 1-2 and 1-2M apply to single auxiliary tank installation for left side, and tables 1-3 and 1-3M are for single tank installedon right side

Rev.3

1-1

T N I R P N O R I O T F C 0 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O Section 1


SUPPLEMENT

IBHT-412-FMS-4.1

Table 1-1.

Fuel loading with both auxiliary tanks (English) (Sheet 1 of 2) TWO 82 U.S. GALLON AUXILIARY FUEL TANKS JET B OR JP-4 (6.5 LB/U.S. GALLON)

QUANTITY (U.S. GAL) 10 20 30 40 50 60 70 *72.6 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 370 374.7

WEIGHT (POUNDS)

65 130 195 260 325 390 455 472 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2405 2435.6

LONGITUDINAL CG MOMENT (INCHES) (IN-LB) 143.6 143.6 140.2 134.8 131.6 129.4 127.9 127.6 128.3 130.4 133.6 136.2 138.2 139.8 141.0 142.0 143.0 143.7 144.5 145.1 145.8 146.3 146.9 147.1 147.7 148.1 148.4 148.7 149.0 149.3 149.6 149.9 150.1 150.4 150.5 150.7 150.9 151.0 151.2

LATERAL MOMENT IS ZERO

9334 18668 27348 35055 42771 50480 58186 60227 66706 76284 86840 93783 107796 118131 128310 138450 148720 158789 169065 179199 189540 199700 210067 219915 230412 240663 250796 260969 271180 281431 291720 302049 312208 322608 332605 342843 353106 363155 368263

* Critical fuel amount for most forward cg condition. I NOTE: All data above representsusable fuel on board (basic and auxiliary),based on nominal density at 15°C. 1-2

Rev. 3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SUPPLEMENT


Section 1

BHT-412-FMS-4.1I

Table 1-1.

Fuel loading with both auxiliary tanks (English) (Sheet 2 of 2) U.S. GALLON AUXILIARY FUEL TANKS JET A, A-1, OR JP-5 (6.8 LB/U.S. GALLON)

TWO 82

QUANTITY (U.S. GAL)

WEIGHT (POUNDS)

LONGITUDINAL CG MOMENT (INCHES) (IN-LB)


10 68 143.6 9765 20 136 143.6 19530 30 204 140.2 28601 40 272 134.8 36666 50 340 131.6 44744 60 129.4 408 52795 70 476 127.9 60880 '72.6 494 127.6 62994 80 544 128.3 69795 90 612 130.4 79805 100 680 133.6 90848 110 748 136.2 101878 120 138.2 816 112771 130 884 139.8 123583 140 952 141.0 134232 150 1020 142.0 144840 160 1088 143.0 155584 170 1156 143.7 166117 180 1224 144.5 176868 190 1292 145.1 187469 200 1360 145.8 198288 210 1428 146.3 208916 220 1496 146.9 219762 230 1564 147.1 230064 240 1632 147.7 241046 250 1700 148.1 251770 260 1768 148.4 262371 270 1836 148.7 273013 280 1904 149.0 283696 290 1972 149.3 294420 300 2040 149.6 305184 310 2108 149.9 315989 320 2176 150.1 326618 330 2244 150.4 337498 340 2312 150.5 347956 350 2380 150.7 358666 360 2448 150.9 369403 370 2516 151.0 379916 374.7 2548.0 151.2 385258 * Critical fuel amount for most forward cg condition. NOTE: All data above representsusable fuel on board (basicand auxiliary), basedon nominal density at I 15°C. Rev. 3

1-3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT MANUFACTURER'S DATA

Section 1

SUPPLEMENT

IBHT-412-FMS-4.1

Table 1-1M. Fuel loading with both auxiliary tanks (Metric) (Sheet 1 of 2) TWO 310 LITER AUXILIARY FUEL TANKS JET B OR JP-4 (0.779kg/liter)

QUANTITY (liters)

WEIGHT (kg)

40 80 120 160 200 240 *274.8 280 320 360 400 440 480 520 560 600 640 680 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1400 1418.4

31.2 62.3 93.5 124.6 155.8 187.0 214.1 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 498.6 529.7 560.9 592.0 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 872.5 903.6 934.8 966.0 997.1 1028.3 1059.4 1090.6 1104.9

LONGITUDINAL MOMENT CG (kgemm/100) (mm) 3647 3647 3541 3399 3322 3272 3241 3246 3277 3360 3432 3493 3538 3571 3602 3627 3650 3668 3688 3703 3719 3731 3744 3754 3764 3774 3782 3790 3797 3805 3810 3815 3823 3828 3833 3835 3840


1137.9 2272.1 3310.8 4235.2 5175.7 6118.6 6939.0 7079.5 8169.6 9421.4 10694.1 11974.0 13228.6 14466.1 15711.9 16952.6 18198.9 19429.4 20686.0 21921.8 23176.8 24415.7 25665.1 26904.9 28147.2 29399.5 30641.8 31885.3 33128.8 34382.0 35615.9 36852.9 38119.1 39363.3 40606.8 41824.5 42428.2

* Critical fuel amount for most forward cg condition. I NOTE: All data above representsusablefuel on board (basicand auxiliary), basedon nominal density at 15°C.

1-4

Rev. 3

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT MANUFACTURER'S DATA

SUPPLEMENT

Section 1

BHT-412-FMS-4.1 2 of 2) (Sheet (Metric) tanks auxiliary both with loading Fuel Table 1-1M. TWO 310 LITER AUXILIARY FUEL TANKS

JET A, A-1, OR JP-5 (0.815kg/liter)

QUANTITY (liters)

40 80

120

WEIGHT (kg)

LONGITUDINAL MOMENT CG (kg.mm/100) (mm)

32.6 65.2

3647 3647

1188.9 2377.8

97.8

3541

3463.1

160 200 240 '274.8 280 320 360 400 440 480 520 560 600 640 680 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1400

130.4 163.0 195.6 224.0 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 554.2 586.8 619.4 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 912.8 945.4 978.0 1010.6 1043.2 1075.8 1108.4 1141.0

3399 3322 3272 3241 3246 3277 3360 3432 3493 3538 3571 3602 3627 3650 3668 3688 3703 3719 3731 3744 3754 3764 3774 3782 3790 3797 3805 3810 3815 3823 3828 3833 3835

4432.3 5414.9 6400.0 7259.8 7407.4 8546.4 9858.2 11188.3 12525.9 13840.7 15133.9 16439.5 17736.0 19038.4 20328.1 21641.2 22936.4 24247.9 25542.4 26852.0 28147.5 29449.5 30758.1 32056.2 33359.6 34659.0 35972.5 37261.8 38554.4 39881.5 41181.6 42485.0 43757.4

1418.4

1156.0

3840

44390.4


* Critical fuel amount for most forward cg condition. NOTE: All data above representsusablefuel on board (basic and auxiliary),basedon nominal density at | 15°C.

Rev.3

1-5



Section 1

BHT-412-FMS-4.1

Table 1-2.

Fuel loading with left auxiliary tank (English) (Sheet 1 of 2)

ONE 82 U.S. GALLON AUXILIARY JET B OR JP-4

QUANTITY (U.S. GAL)

WEIGHT (POUNDS)

(6.5

FUEL TANK ON LEFT

LB/U.S. GALLON)

LONGITUDINAL CG MOMENT (INCHES) (IN-LB)

LEFT LATERAL CG MOMENT (INCHES) (IN-LB)

10 20 30 40 50 60 70 *72.6 80 90

65 130 195 260 325 390 455 472 520 585

143.6 143.6 140.2 134.8 131.6 129.4 127.9 127.6 128.3 131.5

9334 18668 27348 35055 42771 50480 58186 60227 66706 76928

0 0 0 0 0 0 0 0 0 -0.5

0 0 0 0 0 0 0 0 0 -293

100 110 120 130

650 715 780 845

134.6 136.8 138.6 140.2

87490 97812 108108 118469

-1.0 -1.8 -2.5 -3.2

-650 -1287 -1950 -2704

140

910

141.7

128947

-3.9

150

975

143.0

139425

160

1040

144.2

149968

-4.9

-5096

145.2

160446

-5.3

-5857

180 190

1170 1235

146.1 146.9

170937 181422

-5.7 -6.1

-6669 -7534

200 210 220 230 240 250 260

1300 1365 1430 1495 1560 1625 1690

147.7 148.3 148.8 149.4 149.8 150.3 150.6

192010 202430 212784 223353 233688 244238 254514

-6.3 -6.6 -6.8 -7.0 -7.2 -7.4 -7.6

-8190 -9009 -9724 -10465 -11232 -12025 -12844

270 280 290 293.0

1755 1820 1885 1904.5

151.0 151.3 151.6 151.7

265006 275366 285766 288913

-7.8 -7.9 -8.0 -8.0

-13689 -14378 -15080 -15236

170

1105

-4.4

-3549

-4290

* Critical fuel amount for most forward cg condition.

| NOTE: All data above representsusable fuel on board (basicand auxiliary). basedon nominal density at 15°C.

1-6

Rev. 3


Section 1


BHT-412-FMS-4.1

Table 1-2.

Fuel loading with left auxiliary tank (English) (Sheet 2 of 2)

ONE 82 U.S. GALLON AUXILIARY JET A. A-1,

QUANTITY (U.S. GAL) 10 20 30 40 50 60 70 *72.6 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270 280 290 293.0

WEIGHT (POUNDS) 68 136 204 272 340 408 476 494 544 612 680 748 816 884 952 1020 1088 1156 1224 1292 1360 1428 1496 1564 1632 1700 1768 1836 1904 1972 1992.4

OR JP-5

FUEL TANK ON LEFT

(6.8 LB/U.S. GALLON)

LONGITUDINAL CG MOMENT (INCHES) (IN-LB) 143.6 143.6 140.2 134.8 131.6 129.4 127.9 127.6 128.3 131.5 134.6 136.8 138.6 140.2 141.7 143.0 144.2 145.2 146.1 146.9 147.7 148.3 148.8 149.4 149.8 150.3 150.6 151.0 151.3 151.6 151.7

9765 19530 28601 3666 44744 52795 60880 62994 69795 80478 91528 102326 113098 123937 134898 145860 156890 167851 178826 189795 200872 211772 222605 233662 244474 255510 266261 277236 288075 298955 302247

LEFT LATERAL MOMENT CG (IN-LB) (INCHES) 0 0 0 0 0 0 0 0 0 -0.5 -1.0 -1.8 -2.5 -3.2 -3.9 -4.4 -4.9 -5.3 -5.7 -6.1 -6.3 -6.6 -6.8 -7.0 -7.2 -7.4 -7.6 -7.8 -7.9 -8.0 -8.0

0 0 0 0 0 0 0 0 0 -306 -608 -1346 -2040 -2829 -3713 -4488 -5331 -6127 -6977 -7881 -8568 -9426 -10173 -10948 -11750 -12580 -13437 -14321 -15042 -15776 -15939


NOTE: All data above representsusable fuel on board (basic and auxiliary),based on nominal density at I 15°C.

Rev. 3

1-7

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section 1

SUPPLEMENT


I BHT-412-FMS-4.1

Table 1-2M.

Fuel loading with left auxiliary tank (Metric) (Sheet 1 of 2) FUEL TANK ON LEFT

ONE 310 LITER AUXILIARY JET B, OR JP-4

QUANTITY (liters) 40 80 120 160 200 240 *274.8 280 320 360 400 440 480 520 560 600 640 680 720 760 800 840 880 920 960 1000 1040 1080 1109.1

(0.779kg/liter)

WEIGHT (kg)

LONGITUDINAL CG MOMENT (kgmm/100) (mm)

31.2 62.3 93.5 124.6 155.8 187.0 214.1 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 498.6 529.7 560.9 592.0 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 864.0

3647 3647 3541 3399 3322 3272 3241 3246 3277 3381 3452 3505 3548 3589 3627 3658 3686 3711 3734 3752 3769 3785 3797 3810 3820 3830 3840 3848 3853

1137.9 2272.1 3310.8 4235.2 5175.7 6118.6 6939.0 7079.5 8169.6 9480.3 10756.4 12015.1 13266.0 14539.0 15821.0 17097.5 18378.4 19657.2 20944.0 22211.8 23488.4 24769.0 26028.4 27306.3 28566.0 29835.7 31111.7 32373.2 33289.9

LEFT LATERAL CG MOMENT (mm) (kg*mm/100) 0 0 0 0 0 0 0 0 0 -18 -38 -56 -76 -94 -109 -122 -135 -145 -155 -160 -168 -174 -180 -186 -191 -195 -198 -201 -203

0 0 0 0 0 0 0 0 0 -50.5 -118.4 -192.0 -284.2 -380.8 -475.5 -570.2 -673.1 -768.1 -869.4 -947.2 -1047.0 -1138.7 -1233.9 -1333.1 -1428.3 -1519.1 -1604.2 -1691.0 -1753.9


I NOTE: All data above representsusable fuel on board (basicand auxiliary), basedon nominal density at 15°C.

1-8

Rev. 3



Section 1

BHT-412-FMS-4.1I

Table 1-2M.

Fuel loading with left auxiliary tank (Metric) (Sheet 2 of 2)

ONE 310 LITER AUXILIARY JET A, A-1,

QUANTITY (liters) 40 80 120 160 200 240 274.8 280 320 360 400 440 480 520 560 600 640 680 720 760 800 840 880 920 960 1000 1040 1080 1109.1

WEIGHT (kg)

32.6 65.2 97.8 130.4 163.0 195.6 224.0 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 554.2 586.8 619.4 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 903.9

OR JP-5

FUEL TANK ON LEFT (0.815kg/liter)

LONGITUDINAL CG MOMENT (mm) (kg-mm/100)

3647 3647 3541 3399 3322 3272 3241 3246 3277 3381 3452 3505 3548 3589 3627 3658 3686 3711 3734 3752 3769 3785 3797 3810 3820 3830 3840 3848 3853

1188.9 2377.8 3463.1 4432.3 5414.9 6400.0 7259.8 7407.4 8546.4 9919.9 11253.5 12568.9 13879.8 15210.2 16553.6 17887.6 19226.2 20566.4 21911.1 23239.9 24573.9 25912.1 27232.1 28567.4 29887.7 31214.5 32547.8 33870.1 34827.3

LEFT LATERAL CG MOMENT (mm) (kgomm/100)

0 0 0 0 0 0 0 0 0 -18 -38 -56 -76 -94 -109 -122 -135 -145 -155 -160 -168 -174 -180 -186 -191 -195 -198 -201 -203

0 0 0 0 0 0 0 0 0 -52.8 -123.9 -200.8 -297.3 -398.4 -497.5 -596.6 -704.2 -803.6 -909.5 -991.0 -1095.4 -1191.2 -1291.0 -1394.6 -1494.4 -1589.3 -1678.2 -1769.2 -1834.9


NOTE: All data above representsusable fuel on board (basic and auxiliary), basedon nominal density at I 150C.

Rev. 3

1-9



SUPPLEMENT

IBHT-412-FMS4.1 Table 1-3.

Fuel loading with right auxiliary tank (English) (Sheet 1 of 2)

ONE 82 U.S. GALLON AUXILIARY JET B, OR JP-4

QUANTITY (U.S. GAL) 10 20 30 40 50 60 70 '72.6 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 2410 250 260 270 280 290 293.0

WEIGHT (POUNDS) 65 130 195 260 325 390 455 472 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1365 1430 1495 15 1625 1690 1755 1820 1885 1904.5

(6.5

FUEL TANK ON RIGHT

LB/U.S. GALLON)


9334 18668 27348 35056 42771 50480 58186 60227 66706 76928 87490 97812 108108 118469 128947 139425 149968 160446 170937 181422 192010 202430 212784 223353 233688 244238 254514 265006 275366 285766 288913

RIGHT LATERAL CG MOMENT (INCHES) (IN-LB) 0 0 0 0 0 0 0 0 0 0.5 1.0 1.8 2.5 3.2 3.9 4.4 4.9 5.3 5.7 6.1 6.3 6.6 6.8 7.0 7.2 7.4 7.6 7.8 7.9 8.0 8.0

0 0 0 0 0 0 0 0 0 293 650 1287 1950 2704 3549 4290 5096 5857 6669 7534 8190 9009 9724 10465 11232 12025 12844 13689 14378 15080 15236

*Critical fuel amount for most forward cg condition.

I NOTE: All data representsusable fuel on board (basicand auxiliary), basedon nominaldensity at 15C.

1-10

Rev. 3



Section 1

BHT-412-FMS-4.1

Table 1-3.

Fuel loading with right auxiliary tank (English) (Sheet 2 of 2)

ONE 82 U.S. GALLON AUXILIARY JET A, A-1,

QUANTITY (U.S. GAL) 10 20 30 40 50 60 70 '72.6 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270 280 290 293.0

WEIGHT (POUNDS) 68 136 204 272 340 408 476 494 544 612 680 748 816 884 952 1020 1088 1156 1224 1292 1360 1428 1496 1564 1632 1700 1768 1836 1904 1972 1992.4

OR JP-5

(6.8

FUEL TANK ON RIGHT

LB/U.S. GALLON)


9765 19530 28601 36666 44744 52795 60880 62994 69795 80478 91528 102326 113098 123937 134898 145860 156890 167851 178826 189795 200872 211772 222605 233662 244474 255510 266261 277236 288075 298955 302247

RIGHT LATERAL CG MOMENT (INCHES) (IN-LB) 0 0 0 0 0 0 0 0 0 0.5 1.0 1.8 2.5 3.2 3.9 4.4 4.9 5.3 5.7 6.1 6.3 6.6 6.8 7.0 7.2 7.4 7.6 7.8 7.9 8.0 8.0

0 0 0 0 0 0 0 0 0 306 608 1346 2040 2829 3713 4488 5331 6127 6977 7881 8568 9425 10173 10948 11750 12580 13437 14321 15042 15776 15939


NOTE: All data above representsusable fuel on board (basic and auxiliary), basedon nominaldensity at I 150C.

Rev. 3

1-11



SUPPLEMENT

IBHT-412-FMS-4.1 Table 1-3M.

Fuel loading with right auxiliary tank (Metric) (Sheet 1 of 2)

ONE 310

LITER AUXILIARY

JET B. OR JP-4

QUANTITY (liters) 40 80 120 160 200 240 '274.8 280 320 360 400 440 480 520 560 600 640 680 720 760 800 840 880 920 960 1000 1040 1080 1109.1

WEIGHT (kg)

31.2 62.3 93.5 124.6 155.8 187.0 214.1 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 498.6 529.7 560.9 592.0 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 864.0

FUEL TANK ON RIGHT

(0.779 kg/liter)

LONGITUDINAL CG MOMENT (mm) (kgemm/100)

3647 3647 3541 3399 3322 3272 3241 3246 3277 3381 3452 3505 3548 3589 3627 3658 3686 3711 3734 3752 3769 3785 3797 3810 3820 3830 3840 3848 3853

1137.9 2272.1 3310.8 4235.2 5175.7 6118.6 6939.0 7079.5 8169.6 9480.3 10756.4 12015.1 13266.0 14539.0 15821.0 17097.5 18378.4 19657.2 20944.0 22211.8 23488.4 24769.0 26028.4 27306.3 28566.0 29835.7 31111.7 32373.2 33289.9

RIGHT LATERAL CG MOMENT (mm) (kg-mm/100) 0 0 0 0 0 0 0 0 0 18 38 56 76 94 109 122 135 145 155 160 168 174 180 186 191 195 198 201 203

0 0 0 0 0 0 0 0 0 50.5 118.4 192.0 284.2 380.8 475.5 570.2 673.1 768.1 869.4 947.2 1047.0 1138.7 1233.9 1333.1 1428.3 1519.1 1604.2 1691.0 1753.9


I NOTE: All data above representsusable fuel on board (basic and auxiliary),basedon nominaldensity at 15°C.

1-12

Rev. 3

0



Section 1

BHT-412-FMS-4.11

Table 1-3M.

Fuel loading with right auxiliary tank (Metric) (Sheet 2 of 2)

ONE 310 LITER AUXILIARY JET A, A-1.

QUANTITY (liters) 40 80 120 160 200 240 *274.8 280 320 360 400 440 480 520 560 600 640 680 720 760 800 840 880 920 960 1000 1040 1080 1109.1

WEIGHT (kg)

32.6 65.2 97.8 130.4 163.0 195.6 224.0 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 554.2 586.8 619.4 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 903.9

OR JP-5

FUEL TANK ON RIGHT (0.815kg/liter)

LONGITUDINAL CG MOMENT (mm) (kgemm/100)

3647 3647 3541 3399 3322 3272 3241 3246 3277 3381 3452 3505 3548 3589 3627 3658 3686 3711 3734 3752 3769 3785 3797 3810 3820 3830 3840 3848 3853

1188.9 2377.8 3463.1 4432.3 5414.9 6400.0 7259.8 7407.4 8546.4 9919.9 11253.5 12568.9 13879.8 15210.2 16553.6 17887.6 19226.2 20566.4 21911.1 23239.9 24573.9 25912.1 27232.1 28567.4 29887.7 31214.5 32547.8 33870.1 34827.3

RIGHT LATERAL CG MOMENT (mm) (kg-mm/100) 0 0 0 0 0 0 0 0 0 18 38 56 76 94 109 122 135 145 155 160 168 174 180 186 191 195 198 201 203

0 0 0 0 0 0 0 0 0 52.8 123.9 200.8 297.3 398.4 497.5 596.6 704.2 803.6 909.5 991.0 1095.4 1191.2 1291.0 1394.6 1494.4 1589.3 1678.2 1769.2 1834.9


NOTE: All data above representsusable fuel on board (basic and auxiliary), basedon nominal density at I 15°C.

Rev. 3

1-13/1-14

T N I BELL MODELR 412 P N O R I ROTORCRAFT O T F C FLIGHTMANUAL U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-5.1

33001 THROUGH 33107

SUPPLEMENTFOR ALTERNATESTATICAIR VALVE (INSTRUMENTPANELMOUNTED) (412-704-123) CERTIFIED

14 FEBRUARY 1990

This supplement shall be attached to the Model 412 Flight Manual (S/N 33001 through 33107) when the 412-704-123 Alternate Static Air Valve has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and PerformanceData not contained in this supplement, consult the basic Flight Manual.


BHT-412-FMS-5.1


LOG OF REVISIONS

Original ....

0 ....

14Feb90

LOG OF PAGES

Page

Revision No.

Title ....................... A......................... i/ii ........................ 1-4 .......................

Page

Revision No.

0 0 0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT.WORTH, TX 76193-0170

NOTE: Revisedtext is indicated by a black vertical line. Insert latest revision pages. Disposeof superseded pages.

A

Useor disclosureof data containedon this pageis subject to the restriction on the title pageof this document.



INTRODUCTION

BHT-412-FMS-5.1

INTRODUCTION

The Alternate Static Air Valve provides an alternate static source (cabin air). The kit consists of a toggle valve, static lines, and associated hardware.

Useor disclosureof data containedonthis pageis subject to the restriction on the title pageof this document.

i/ii

T Section 1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT


Sections 1 & 2

BHT-412-FMS-5.1

PLACARDSAND DECALS

PRESTARTCHECK

STATICSOURCEswitch - PRI.


1

T N I 0 Section 3 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I Section 4 R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL

Sections 3 & 4

BHT-412-FMS-5.1


STATICPORTOBSTRUCTION INDICATION:

Erratic readings from the airspeed indicator, vertical speed indicator, and altimeter when operating helicopter in rain with STATICSOURCE switch in PRIposition.

PROCEDURE: 1. 2. 3.

Windows and vents - Close. HEATERswitch - OFF. STATICSOURCEswitch - ALTN.

NOTE:

This procedure selects an alternate static source Icabin air) for pilot side instruments only.

No change from basic Flight Manual

Useor disclosureof data containedon this pageis subject to the restrictionon the title pageof this document.

T N I * Section 1 R P N O R I O T F C * TSection 2 U O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT MANUFACTURER'SDATA

Section 1 & 2

BHT-412-FMS-5.1

MANUFACTURER'SDATA


MANUFACTURER'S DATA

I

INSTRUMENTPANELAND CONSOLES

ou

I AI

PILOTPANEL

Figure 2-1.

Instrument Panel


3

T N I R P N O R I O T F C U T O OD N R E P E E R Section 3 SCR N G O N I R W O FSection 4VIE LY N O Sections 2, 3, & 4

412 ROTORCRAFT MANUFACTURER'SDATA

BHT-412-FMS-5.1

PITOT-STATICSYSTEM

An alternate static port is located inside the cockpit on the pilot instrument panel on back of the STATICSOURCE switch (figure 2-1). Under normal conditions, the switch should be placed in the PRI position. This position selects the static ports located forward of the crew doors as well as the roof mounted static port (if installed). If erratic readings are seen on the airspeed indicator, altimeter, and vertical speed indicators, obstruction of the outside static ports is a possible cause. If this occurs, the STATIC SOURCE switch should be placed in the ALTN position. This position selects the alternate static air source (cabin air) and at the same time. shuts off the outside static air source for the pilot side only.

MANUFACTURER'SDATA

I


MANUFACTURER'SDATA


0

4



SUPPLEMENTFOR FLIGHT DIRECTOR (412-706-111)

CERTIFIED FEBRUARY13, 1981

Manual when the 412-706-111 installed.

Flight Director has been

The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter E i .*

tl

I

A Subsidiary of lexfon Inc POST OFFICE BOX 482 · FORT WORTH, TEXAS 76101

COPYRIGHT1989 BELL:HELICOPTER TEXTRON INC. ALL RIGHTS RESERVED

13 13

FEBRUARY 1981

FEBRU ARY 198 REISSUED8 MAY 1989



BHT-412-FMS-6

LOGOF REVISIONS

Original .... Reissued....

0 .... 13 Feb81 0 ... .8 May 89

LOGOF PAGES

FLIGHTMANUAL

Page

Title - A................... i/ii......................... 1 -5......................

Revision No.

MANUFACTURER'SDATA

Page

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

0 0 0

Revision No. 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATION DIRECTORATE AIRCRAFTCERTIFICATIONSERVICE DEPARTMENT OF TRANSPORTATION SOUTHWESTREGION,FORTWORTH.TEXAS


A

Useor disclosure of data containedon this pageissubjectto the restrictionon the title pageof this document.




INTRODUCTION

The Flight Director assists the flight crew in control and

The Flight Director can be coupled to the Automatic Flight Control System (AFCS) for fully automatic (hands off) flight path control.

Use or disclosure of data contained on this page is subject to

the restrictionon the title pageof this document.

i/ii



Section 1

BHT-412-FMS-6

TYPE OF OPERATION

The Flight Director may be used during VFR or IFRnonicing conditions.

FLIGHT DIRECTORLIMITATIONS

NOTE

Flight Director bars are repeated on copilot ADI.

WEIGHT/CG

LIMITATIONS

During VOR approaches, except for VORs collocated at the airport, the Flight Director shall not be coupled in VOR APR mode prior to VOR station passage inbound.

Actual weight change shall be determined after the Flight Director is installed and ballast readjusted, if necessary, to return empty weight CGwithin allowable limits.

Maximum approach gradient is 5 degrees.

AIRSPEED LIMITATIONS

Flight Director modes may not be commanded from the copilot position.

Minimum airspeed for coupled operation of Flight Director is 60 KIAS.

Useor disclosure of datacontainedon this pageissubjectto the restrictionon the title pageof this document.

1

T N I R Section 2 P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section 2


BHT-412-FMS-6

PRESTART CHECK PILOT INSTR LT operation).

Flight director VS button - Depress; check VS light illuminates.

knob

-

OFF (day

ADI pitch command bar - Check centered.

AUX SYS PITOT and STATIC switches NORMAL.

Cyclic ATTD TRIM switch - Move aft; check ADI pitch command bar moves up and pitch API moves up; check cyclic stick moves aft in approximately 2 seconds.

SYSTEMS CHECKS

FLIGHT DIRECTOR CHECK

Cyclic NAV STBY button Depress momentarily and release. Check HDG.VS, and CPL lights extinguish. SBY light illuminates; check all APIs center, pitch and roll command bars retract from view.

FORCE TRIM switch - ON.

HP1 and HP2 buttons - ON.

Flight director SBY button - Depress and hold; check DCPL light and all mode selector and helipilot controller lights illuminate and FD fail flag on ADI appears; then release button. HSI heading marker - Set to aircraft heading. Flight director HDG button - Depress; check HDG and CPL lights illuminate, SBY light extinguishes.

ADI roll command bar - Check centered.

HSI heading marker - Move right; check ADI roll command bar moves right and roll actuator position indicator (APII moves right; check cyclic stick moves right in approximately 2 seconds.

HSI heading marker heading.

2


Reset to aircraft

HSI heading marker - Set to takeoff heading.

NOTE

For prolonged ground operation, AFCS shall be operated in SAS mode only.

TAKEOFF

SAS/ATT button - ATT or SAS as desired during hover and takeoff; ATT prior to entering Instrument Meteorological Conditions (IMC). NOTE

It is recommended that the cyclic FORCE TRIM release button be depressed before liftoff to trim actuators to center positions.

Useor disclosureof datacontainedonthis pageissubjectto the restrictionon thetitle pageof this document.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S Section 3 N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

412 ROTORCRAFT FUGHT MANUAL

Sections2 &3

BHT-412-FMS-6

Cyclic FORCETRIM release button - Depress and hold until desired climbout attitude is attained, then release.

Collective pitch power setting.

Airspeed - Adjust to desired climb speed.

Flight director - Select modes as desired after reaching 60 KIAS.

Flight director - Select modes as desired.

IN-FLIGHT OPERATION

LANDING

Flight director - Select modes as desired.

Cyclic NAV STBY button - Depress at or above 60 KIAS to decouple flight director.

DESCENT AND APPROACH

AFCS - SAS or ATT mode as desired.

Flight director - Select mode for type of approach to be flown.

NOTE

Collective pitch - Adjust to maintain desired approach speed.

It is recommended that the cyclic FORCE TRIM release button be depressed before touchdown to trim actuators to center positions.

GO-AROUND

Collective GO-AROUND button - Depressat Missed Approach Point.

Table 3-1.

CAUTION LIGHT WORDING DCPL

Adjust to desired climb

For prolonged ground operation, AFCS shall be operated in SAS mode only.

Caution lights

FAULT CONDITION

Flight director not coupled.

Use or disclosure of data contained on this page is subject to the restriction on the title page of this document.

CORRECTIVE ACTION

Ensure that HP1 and HP2 are engaged in ATT mode. Engageflight director modes as desired. DepressCPL button if CPL light not illuminated.




BHT-412-FMS-6

Table 3-2.

FLAG WORDING ATT

FLAG LOCATION ADI

Warning flags

FAULT CONDITION

Vertical gyro not turning at full speed/attitude information unreliable. (Flight director may or may not decouple.)

CORRECTIVE ACTION

If only pilot ATT flag is displayed, pilot shall monitor standby attitude indicatorand flight director.

If only the copilot ATT flag is displayed, the copilot shall monitor the standby attitude indicator. Flight director is functional and flight director indicationsrepeatedon copilots ADI are valid.

Check PILOTand CPLTATTD SYScircuit breakers- in. If flag does not retract, continue flight in ATT or SAS mode.

FD

ADI

Flight director not coupled due to flight director failure. (Pitch and roll command bars may or may not retract from view.)

Check FLT DIR(AC and DC) circuit breakers in. Check desired mode(s) engaged. If flag does not retract from view, continue flight in ATT or SAS mode. (ATT mode for IMC conditions.)

GS (No legend)

HSI

Glideslope signal unreliable. (ADI pitch command bar retracts from view. AFCS holds pitch attitude present at time of signal failure.)

Continue flight in any mode, using cyclic ATTD TRIM switch and collective to maintain airspeed and glidepath. (Monitor raw glideslope data.)

NAV (No legend)

HSI

VOR or localizer signal unreliable. (ADI roll command bar retracts from view. HSI course deviation bar and No. 1 bearing pointer unreliable. AFCS holds roll level attitude.)

Check NAV 1 (DC)circuit breaker in. Check VHF NAV 1 tuned properly. signal identified. Continue flight in HDG or ATT mode.

OFF

HSI

Directional gyro failure or HSI failure. Headinginformation unreliable. (AFCS holds roll level attitude.)

Check PILOTand CPLTHSI and GYRO CMPS (AC) circuit breakers in. Continue flight in any mode except HDG.


T N I R P N O R I O T F C U T D O Section 4 O N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHTMANUAL


Sections 3 & 4

BHT-412-FMS-6

FLIGHT DIRECTOR FAILS TO COUPLE HP1 and HP2 buttons - ON. SAS/ATT button - ATT. CPL button illuminated.

Depress if CPL light not

PITOT-STATIC SYSTEM MALFUNCTION

In the event of an apparent malfunction of the copilot altimeter, vertical speed indicator, and/or airspeed indicator, proceed as follows:

AUX SYS PITOT or STATIC switch (as applicable) - OFF.

NOTE

The Flight Director will not couple if HP1 or HP2 is inoperative.

Flight director modes.

-

Disengage

vertical



5

T N I Section 1 R P N O R I O T F C U T Section 2 O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Sectons 1 & 2


SUPPLEMENT

BHT-412-FMS-6

MANUFACTURER'SDATA

No changefrom basic Flight Manual

MANUFACTURER'SDATA

FLIGHT DIRECTOR(412-706-1 11)

Capture and track an ILS localizer and glideslope (ILS).

The flight director is designed for use as a workload reliever to assist the pilot in control and navigation of the helicopter. The flight director has nine modes of operation, any of which may be coupled to the helipilot system for fully automatic hands off flight path control. When decoupled from the helipilots. the flight director provides automatic flight path computation and visual pitch and roll command indications to direct the pilot in maneuvering the helicopter to maintain the selected flight path. When the flight director is coupled in the appropriate modes, the automatic flight control system will maneuver the helicopter to perform the following functions:

Capture and track a localizer back course (BC).

Maintain a constant pressure altitude (ALT).

Maintain a constant indicted airspeed IlAS).

Maintain a constant vertical speed climb or descent (VS). Turn to and maintain a selected magnetic heading (HDG). Capture and track a selected VOR radial (NAV or VOR APR).

6

Initiate a missed approach (go-around) climbout (GA).

The flight director computer analyzes vertical and lateral flight and navigational data to generate pitch and roll steering commands which are displayed visually on the attitude director indicator (ADI). The vertical channel combines pitch attitude, airspeed, altitude. vertical speed, and glideslope deviation information to produce computed pitch command signals. The lateral channel combines roll attitude, heading, and course deviation information to produce computed roll command signals. Automatic flight path control is achieved when the pitch and roll commands from the flight director computer are coupled to the helipilot computers. Should a flight or navigation data signal become invalid, the affected pitch or roll channel will revert to attitude hold mode and the respective command bar on the ADI will retract from view. If either helipilot fails or is disengaged, the flight director will decouple automatically.

Jseor disclosureof datacontainedon this pageis subjectto t herestrictionon the title pageof this document.



Section 2

BHT-412-FMS-6

FLIGHT DIRECTOR CONTROLS

MODE SELECTOR

The modeselector (figure 2-1) enablesthe pilot to select the desired flight director mode by depressing the appropriate push-on/push-off button. Selected modes are annunciated by illumination of the respective mode select buttons. The NAV, BC, and VOR APR buttons have two lights each (ARM and CAP) to advise the pilot of the status of the flight director computer. The ILS button legends (ARM and GS) advise when the computer has armed or captured the glideslope. Depressing the SBY button disengagesall modes and tests the illumination of all mode selector lights. Turning the PILOTINSTR LT knob clockwise dims all mode selector lights for night operation.

CPL button. When decoupled, the flight director will continue functioning in the selected mode, providing visual pitch and roll commands to the pilot on the attitude director indicator (ADI). Once depressed, the CPL button must be depressed again to recouple any flight director mode. FORCE TRIM SWITCHES

The FORCETRIM switch on the pedestal must be ON anytime the flight director is coupled for automatic flight control.

The cyclic mounted FORCETRIM release button can be depressed to allow the pilot to reposition the cyclic control and pedals manually for large scale pitch, roll, and yaw corrections. Upon depressing the button, the rotary trim actuators are de-energeized. the flight director modes are decoupled momentarily, the helipilot pitch, roll, and yaw linear actuators return to center positions, and the helipilot computers are placed in a fast follow-up mode to track flight control positions. Upon releasing the button, the helipilots and flight director will resume functioning in the preselected modes. If previously decoupled in the ATT mode, the helipilots will maintain the attitude existing at the time the button is released. ATTITUDE TRIM SWITCH

Figure 2-1.

Flight director mode selector

COUPLE BUTTON

The CPL button is a push-on/push-off button located on the helipilot control panel on the pedestal. When both helipilots (HP1 and HP2) are engaged in the attitude retention mode (ATT), selecting any valid flight director mode will couple the flight director to the helipilots automatically, as indicated by illumination of the ON legend of the CPL button. The pilot may decouple the flight director by depressing the

The cyclic mounted ATTD TRIM switch can be moved fore and aft to adjust pitch attitude and laterally to adjust roll attitude during decoupled operation in attitude retention mode. The switch is also used to make small pitch attitude changes when coupled in any mode except ALT and ILS after glideslope capture. Roll attitude can also be adjusted by the ATTD TRIM switch, except when a lateral mode is engaged. (Large attitude changes should be made by depressing the cyclic FORCETRIM release button.)

Useor disclosureof data containedon this pageis subjectto the restrictionon the title pageof this document.

FLIGHTDIRECTORSTANDBY BUTTON

The cyclic mounted NAV STBY button is a remote switch having the same function as the SBY button on the flight director mode selector. Depressing the button disengages all flight

7



BHT-412-FMS-6 director modes, tests the illumination of all motde selector lights, retracts the pitch and re command bars on the ADI, and places the flig il director in a standby status. The pilot must the reselect the modes if continued flight direct operation is desired. GO-AROUND BUTTON

The GO-AROUND button, located on tlhe collective control head, is a remote switi ch having the same function as the GA button eon the flight director mode selector. Depressing tlhe button places the flight director in go-arournd mode and disengages all other modes. In GiA mode the flight director commands a roll lev el attitude and a pitch attitude which will providesa 750 feet-per-minute rate of climb. (The pilot must adjust collective pitch to maintain desireed climb airspeed.)

8

AUXILIARY PITOT-STATIC

SUPPLEMENT

SWITCHES

The AUX SYS PITOT and STATIC switches provide a means for isolating the flight director airspeed and altitude sensors from the copilot pitot-static system in the event of leakage or other system malfunction. When the PITOT switch is OFFthe flight director airspeed sensor is disconnected. rendering the IAS mode inoperative. Placing the STATIC switch in the OFF position disconnects both the airspeed and altitude sensors of the flight director and thereby disables the ALT. IAS, and VS modes. The pilot should disengage the affected vertical modes to prevent undesirable flight control inputs when either switch is in the OFFposition.

of data containedonthis pageis subjectto Useor disclosure the restrictionon the title pageof this document.

T N I R P N ROTORCRAFT O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-7

FLIGHT MANUAL SUPPLEMENT

SINGLE-SPEEDINTERNALHOIST 214-706-003

CERTIFIED OCTOBER2,1981

This supplement shall be attached to the Model 412 or 412EP Flight Manual when the 214-706003 Single Speed Internal Hoist has been installed.

Information contained herein supplements of basic Flight Manual. For information Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter NOTICE COPYRIGHT

Inc of Textron A Subsidiary

®

BELL HELICOPTER INC INC TEXTRON ANDBELLHELICOPTER ADIVISION OFTEXTRON CANADA LTD ALLRIGHTS RESERVED

POSTOFFICEBOX482· FORT WORTH, TEXAS76101

REISSUE -

MAY 11, 1995


FAA APPROVED

NOTICEPAGE

PROPRIETARY RIGHTS NOTICE

Manufacturer's Data portion of this supplement is proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron Inc.

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 76101-0482


BHT-412-FMS-7

LOG OF REVISIONS

Original...........0 Reissued .........

........ October 2, 1981 0 ............ May 8, 1989

Reissued ......... Reissued .........

0 ........ October 8, 1991 0 ........... May 11, 1995

LOGOF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title -

NP .............................

A- B .................................. i/ii ....................................... 1 - 34 ..................................

REVISION NO.

PAGE

35/36....................................

MANUFACTURER'SDATA

0

0 0 0

0

37 -

38.................................

0

NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. A

T N I R P N O R I O T F C U T 0 O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-7

LOGOF APPROVED REVISIONS

Original ........... Reissued .........

0 ......... October 2,1981 ............ May 8, 1989

Reissued ......... Reissued .........

0.........October 8,1991 0 ......... May 11, 1995

APPROVED:

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERALAVIATIONADMINISTRATION FT. WORTH,TX 76193-0170

B

0


BHT-412-FMS-7

GENERAL INFORMATION

The Single Speed Internal Hoist enables cargo and emergency rescue operations in areas where landings cannot be accomplished. The hoist can raise or lower loads up to 600 pounds (272 kilograms). The hoist cable is 250 usable feet (76.2 meters) In length. Each of the four cabin mounting locations allows the hoist to be extended 90 degrees outboard. Caution

lights on each side of the hoist illuminate when the hook reaches 20 feet (6 meters) below the skids during retraction. An electrically actuated cable cutting device allows pilot or hoist operator to sever the cable in an emergency. A manually operated cable cutter, accessible to the hoist operator, may be used to sever the cable if the electrical cable cutter fails.

i/ii


FAA APPROVED

Section

1

1-3. TYPES OF OPERATION

Hoist operation with flight coupled mode is prohibited.

Hoist operations shall be conducted under appropriate operating rules for external loads.

Hoisting or lowering an empty litter in open position (except Stokes litter) is prohibited.

Passenger operations with hoist installed are approved if hoist is stowed and

Refer to appropriate Flight Manual Supplement(s) for additional limitations,

electrical system is deactivated,

Hoist operations are prohibited during instrument

1-4.

meteorological

conditions.

FLIGHT CREW

A crewmember wearing an approved safety

harness in passenger compartment is

required during all phases of hoist operations. Crewmember shall wear protective gloves for guiding cable during operation. The hoist operator shall be familiar with hoist operating procedures and limitations.

1-5.

CONFIGURATION ÛI~~~rmun~only

1-5-A. 1-5-A. REQUIRED REQUIRED EQUIPMENT EQUIPMENT

Hoist cable antichafing guard shall be installed on standard or high skid landing gear (with or without floats) on same side

of helicopter as hoist.

in

director

procedures, and performance data.

1-6. WEIGHT AND CENTER OF GRAVITY Actual weight change shall be determined

after hoist is installed readjusted, if necessary.

and ballast

For maximum gross weight, including hoist load, refer to applicable Flight Manual or BHT-412-FMS-19.1 when Increased Gross Weight and Takeoff Horsepower kit is installed.

Maximum hoist load is 600 pounds (272 kilograms). This is a structural limitation

and does not ensure that longitudinal or lateral CG will remain within approved limits.

Maximum

allowable

hoist

load

varies with gross weight, center of gravity, and hoist location. Refer to appropriate Hoist Loading Schedule. NOTE

The center of gravity of hoist load

1-5-B.

OPTIONAL EQUIPMENT

Fixed passenger step shall not be installed

concurrently with internal hoist.

Retractable passenger steps shall be stowed during hoist operations.

in forward position

is F.S. 82

(2083 mm) and B.L. 60 (1524 mm). The center of gravity of hoist load

in aft position is F.S. 131 (3327 mm) and B.L. 64.4 (1636 mm).


FAA APPROVED

For Longitudinal vs. Lateral CG limits with

internal hoist refer to Internal envelope figure 1-1.

1-7.

hoist CG

Enter appropriate schedule, figures 1-2

AIRSPEED

through 1-5 at gross weight of helicopter

VNE with asymmetrical door configuration is 20 KIAS.

VNE with hinged panels locked open and cargo doors open is 20 KIAS. VNE with hinged panels removed and cargo doors removed or secured open is 60 KIAS.

1-24.

1-25-A. LEFT HOIST INSTALLATIONS INSTALLATIONS

HOIST DUTY CYCLE

HOIST DUTY CYCLE

LIMITATIONS

The hoist is approved for continuous operation with loads not to exceed 600 pounds (272 kilograms).

prior to hoisting. Proceed vertically to intersect with diagonal line representing number of crewmembers on board, top of schedule, or right cutoff line. Proceed

horizontally to left to read maximum allowable hoist load. Intersecting with right cutoff line gives maximum load which does not cause helicopter to exceed gross

weight limitations.

Using Weight empty chart, Section 5 and

left hoist loading schedules ensures that both longitudinal and lateral limits are not

exceeded during first hoist operation. However, for subsequent hoisting, additional precautions must be taken to avoid exceeding forward longitudinal limits. 1-25-A-1.

1-25.

ALLOWABLE

HOIST

LOAD

Select hoist loading schedules (figures 1-2 through 1-5) appropriate for position in which hoist is installed.

NOTE

Hoist loading schedules are based on most adverse loading combinations of pilot, copilot, and hoist operator, each weighing 170 or 200 pounds (77.1 or 90.7

LEFT FORWARD HOIST

LOCATION To continue using maximum allowable

hoist capability: (Refer to figure 1-2

through 1-5)

a. put hoisted load (people or cargo) along side of island, or

b. when hoisted

load is put

immediately forward of island, reduce maximum hoist load to 300 pounds.

kilograms), and on a weight empty CG of 0.3 inches (7.3 mm) to right

WARNING

of centerline prior to adding hoist. If lateral CG is appreciably different or crewmember weights

DO NOT PUT HOISTED LOAD IN FORWARD AREA OF PASSENGER

hoist load shall be computed. For

MAXIMUM

are out of this range, allowable

computation, assume hoist

operator in forward position to be located at F.S. 87 (2210 mm) and

B.L. 40 (1016mm), and in aft position F.S. 125 (3175mm) and

B.L. 40 (1016mm). 2

COMPARTMENT

HOIST

COMPUTED

UNLESS

LOADS

CONFIGURATION. 1-25-A-2.

LEFT AFT HOIST LOCATION

ARE

FOR THAT


BHT-412-FMS-7

To continue using maximum allowable hoist capability: (Refer to figure 1-2 through 1-5) a. put hoisted load along Island or Immediately forward of Island, or b. ensure empty weight CG Is within Area A. Refer to Weight empty chart, Section 5.

1-25-B. RIGHTHOIST INSTALLATIONS

Right lateral limit for hoist operations varies with longitudinal center of gravity of the helicopter. The loading schedules have been modified to account for this variation. a. Starting with appropriate schedule for number of crewmembers on board, enter at gross weight of helicopter prior to hoisting. b. Proceed vertically to Intersect with diagonal line representing helicopter center of gravity prior to hoisting, top of schedule, or right cutoff line. c. Proceed horizontally to left to read maximum allowable hoist load.

When helicopter center of gravity between STA. lines, Interpolate determine CG.

Is to

GIVEN:

Gross Weight - 9,500 Ibs. CG - STA. 135.5 before hoisting

From appropriate 11,600 lb. GW schedule obtain hoist load as follows:

Enter gross weight at 9,500 Ibs. Proceed up GW line to Interpolated STA. 135.5 Proceed left to read hoist load of 210 Ibs. Point .

EXAMPLE2: NORMAL

Determine Hoist Load when hoist is In R/H FWD POSITIONand crew consist of Pilot, Copilot and Hoist Operator. GIVEN:

Gross Weight - 9,500 lbs. CG - STA. 138.5 before hoisting


Enter gross weight at 9,500 lbs. Proceed up GW line to STA. 138.5 Proceed left to read hoist load of 550 lbs. Point .

Intersecting right cutoff line gives maximum load which does not cause helicopter to exceed gross weight limitations or forward longitudinal limits.

1-25-C. RIGHTHOIST

For multiple hoists during a single flight, after each hoist operation enter appropriate schedule at revised gross weight and proceed to new center of gravity to determine maximum allowable hoist load. EXAMPLE1: NORMAL

The dashed line on schedules represents longitudinal center of gravity prior to hoisting which will result In a gross weight center of gravity at Sta. 135.2 and B.L. 4.5 during hoist operations with maximum hoist loads derived using this line. This center of gravity Is the corner of but not In Penalty Region shown In Limitations.

Determine Hoist Load when hoist Is In R/H FWD POISTIONand crew consist of Pilot, Copilot and Hoist Operator.

Hoist loads derived for Normal Operations may be Increased when GW/CG combinations are forward of those

INSTALLATION - PENALTY REGION

OPERATION

3


FAA APPROVED

represented by dashed line. Loads may be increased up to but not greater than those defined by dashed line. However, this procedure will result In operations within Penalty Region. Refer to Section 1, Internal Hoist CG Envelope, for Penalty Region. EXAMPLE 3: PENALTY REGION

Determine Hoist Load when hoist Is In R/H FWD POSITION and crew consist of Pilot, Copilot and Hoist Operator.

For GW vs. CG combinations aft of the CG represented by the dashed line (see Example 2), there Is no Penalty Region.

1-26. WEIGHT EMPTY CHART

The Weight empty chart for internal hoisting operations Is shown In Section 5. Refer to the maintenance manual for additional information. NOTE

GIVEN:

Gross Weight CG -

9,500 lbs.

Allowable hoist load must be computed when weight empty Is not within specified guidelines, shown In Section 5.

STA. 135.5 before hoisting

NOTE

From appropriate 11,600 lb. GW schedule obtain hoist load as previously determined In Example 1 the maximum hoist load for normal operations Is 210 lbs. Point (). To increase hoist load to maximum for condition without exceeding GW/CG limits, proceed up to dashed line and read left to find 435 lbs. Point .

The Penalty Region Is any load greater than Point up to maximun load at Point .

4

Allowable hoist loads must be computed when AUX Fuel kits are Installed.

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O Longitudinal/Lateral C.G. Envelopefor InternalHoistOperations

132 (3353)

135

(3429)

136

(3454)

137

(3480)

138

Z

(3505)

139

(3531)

140

(3556)

141

(3581)

142

(3607)

143

(3632)7

(-178)

-6

-5

(-152)

(-127)

-4

(-102)

-3

-2

-1

0

1

2

3

4

5

6

(-76)

(-51)

(-25)

(0)

(25)

(51)

(76)

(102)

(127)

(152)

LATERAL - INCHES (MILLIMETERS)

(178)

412FS.7.1-1

1

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O 0 0 LB

HOIST LOAD IS ALLOWED AT GROSS WEIGHTS FROM 7133 TO 11000 LBS. 1 PILOT AND HOIST OPERATOR - 6 12 PILOT, COPILOT, AND HOIST OPERATOR - 600 LB HOIST LOAD IS ALLOWED AT GROSS WEIGHTS FROM 7866 TO 11000 LBS. 11000

o

0@ I

I

@I~'

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O LH AFT POSITION

PILOTAND HOISTOPERATOR- 600 LB HOISTLOADIS ALLOWEDATGROSSWEIGHTS FROM7133 TO 11000 LBS. 2 PILOT,COPILOT,AND HOISTOPERATOR- 600 LB HOISTLOADIS ALLOWEDATGROSSWEIGHTSFROM 7866 TO 11000LBS. 11000

600

500

400

HOISTLOADING

300

I w

200

100

0

7(7000

7500

8000

8500

GROSS WEIGHT

9000

9500

10000

PRIOR TO HOIST OPERATION

10500

11000

11500

12000

- POUNDS

412FS-7-1-2-2

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H FWDPOSITION- PILOTAND HOISTOPERATOR

PROHIBITED

11500 11000 10000 10500 9500 9000 8500 GROSS WEIGHT PRIOR TO HOIST OPERATION - POUNDS NOTE: THE DASHED LINE INDICATES A DIFFERENT LONGITUDINAL C.G. FOR EACH GROSS WEIGHT.

7000

7500

8000

12000

412FS-7-1-2-3

m

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H FWD POSITION- PILOT,COPILOTAND HOIST OPERATOR

HOIST LOADING IN THIS AREA

PROHIBITED

7000

7500

8000

8500 9000 9500 10000 10500 GROSS WEIGHT PRIOR TO HOIST OPERATION - POUNDS

11000

11500

12000

NOTE: THE DASHED LINE INDICATESA DIFFERENT LONGITUDINAL C.G. FOR EACH GROSS WEIGHT.

412FS-7-1-2-4

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H AFT POSITION - PILOT AND HOIST OPERATOR

11000

o

7000

7500

8000

10000 10500 8500 9000 9500 GROSS WEIGHT PRIOR TO HOIST OPERATION - POUNDS

11000

11500

12000

NOTE: THE DASHED LINE IS A CONSTANT LONGITUDINAL C.G. FOR ALL GROSS WEIGHTS.

412FS-7-1-2-5

m

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H AFT POSITION - PILOT,COPILOT AND HOIST OPERATOR

500

o

HOIST LOADING IN THIS AREA PROHIBITED

o

7000

7500

8000

8500 9000 9500 10000 10500 11000 GROSS WEIGHT PRIOR TO HOIST OPERATION - POUNDS

INOTE: THE DASHED LINE IS A CONSTANT LONGITUDINAL C.G. FOR ALL GROSS

11500

12000

WEIGHTS.

412FS-7-1-2-6

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O L/HFWD POSITION

1 PILOTAND HOISTOPERATOR- 600 LB HOISTLOAD IS ALLOWEDAT GROSSWEIGHTS FROM 3233.8TO 4989.5 KGS. 2 PILOT,COPILOT,AND HOISTOPERATOR- 600 LB HOISTLOADIS ALLOWEDATGROSSWEIGHTSFROM 3566.5TO 4989 5 KGS. 4989.5

I

250

200

3000

3250

3500

3750 4000 4250 4500 4750 5000 GROSS WEIGHT PRIOR TO HOIST OPERATION - KILOGRAMS

5250

5500

412FS-7-1-3-1

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O L/H AFT POSITION

1 PILOTAND HOISTOPERATOR- 600 LB HOISTLOADIS ALLOWEDATGROSSWEIGHTS FROM 3498.2TO 4989.5 KGS. 2 PILOT COPILOT,AND HOISTOPERATOR- 600 LB HOISTLOADIS ALLOWEDAT GROSSWEIGHTSFROM3830.8 TO 4989 5 KGS 4989 5

250

412FS-7-1-3-2

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H FWDPOSITION- PILOTAND HOIST OPERATOR

4989.5

250

0

a u)

INTHISAREA PROHIBITED

3000

3250

3500

3750

4000

4250

4500

4750

5000

5250

5500

GROSS WEIGHT PRIOR TO HOIST OPERATION - KILOGRAMS

NOTE: THE DASHED LINE INDICATESA DIFFERENT LONGITUDINAL C.G. FOR EACH GROSS WEIGHT.

412FS-7-1-3-3

I R N P O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O c

0

o0

IN THIS AREA

PROHIBITED

3000

3250

3500

3750

GROSS WEIGHT

4000

4250 4500 PRIOR TO HOIST OPERATION

4750

5000

5250

5500

- KILOGRAMS

NOTE:THEDASHEDLINEINDICATES A DIFFERENTLONGITUDINAL C.G.FOREACHGROSSWEIGHTI

412F6-7-1-3-4

I R N P O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H AFT POSITION- PILOTAND HOIST OPERATOR

o

3

HOIST LOADING IN THIS AREA PROHIBITED

50

3000

3250

3500

3750 4000 4250 4500 4750 5000 GROSS WEIGHT PRIOR TO HOIST OPERATION - KILOGRAMS NOTE: THE DASHED LINE IS A CONSTANT LONGITUDINAL C.G. FOR ALL GROSS WEIGHTS.

5250

5500

412FS-7-1-3-6 m

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H AFT POSITION - PILOT, COPILOT AND HOIST OPERATOR

4989.5

200

150

IN

THIS

HOIST LOADING

AREA

PROHIBITE

00

GROSSWEIGHTPRIORTO HOISTOPERATION - KILOGRAMS

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O L/H FWD POSITION

1 PILOTAND HOISTOPERATOR-600 LB HOIST LOADIS ALLOWEDATGROSSWEIGHTSFROM7133 TO 11300LBS. 2 PILOT.COPILOT.AND HOISTOPERATOR- 600 LB HOISTLOADIS ALLOWEDATGROSSWEIGHTSFROM7866 TO 11300LBS. 11300

600

500

INTHIS AREA

400

7000

7500

8000

7000

7500

8000

8500 9000 9500 10000 10500 11000 GROSSWEIGHTPRIORTO HOISTOPERATION - POUNDS 8500

9000

9500

10000

10500

11000

11500

12000

412FS-7-1-4-1

11500

12000


BHT-412-FMS-7

o 0

o

o

Figure 1-4.

Hoist loading schedules 11,900 lb. GW (English) (Sheet 2 of 6)

19


Figure 1-4.

20

FAA APPROVED



0

z

I0

z0 0

Figure 1-4.


21

T N I R P N O R I T IFO C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-7

Figure 1-4.

22

FAA APPROVED



BHT-412-FMS-7

o

o

Figure 1-4.


23


Figure 1-5.

24

FAA APPROVED

Hoist loading schedules 5397 kg. GW (Metric) (Sheet 1 of 6)


Figure 1-5.

BHT-412-FMS-7


25


Figure 1-5.

FAA APPROVED



Figure 1-5.

BHT-412-FMS-7


27


FAA APPROVED 0

Figure 1-5.



Figure 1-5.

BHT-412-FMS-7


29

T N I Section 2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I I R W O E I F V NLY O BHT-412-FMS-7

FAA APPROVED

2-2. FLIGHTPLANNING WARNING

HOIST LOAD CAN CAUSE LONGITUDINAL OR LATERAL CG LIMITS TO BE EXCEEDED.GROSS WEIGHT AND CENTER OF GRAVITY SHALL BE COMPUTED TO ASSURE LOADING WITHIN APPROVED LIMITS.

CAUTION

IF ADDITIONAL LOADS ARE CARRIED DURING HOISTING OPERATIONS, LOADS SHOULD BE PLACED ON SIDE OF HELICOPTER OPPOSITE HOIST POSITION.

Gross weight and CG without hoist load.

Hoist

-

Installed

in desired

position;

check roof and floor stud adapters and locking collars properly secured. Boom actuator - Installed position; all fittings secured.

in proper

AIRCRAFT POSITION switch (on hoist control box, figure 2-1) - Set in proper position. Hook -

Rotates freely on cable.

Cable - Check proper routing through guide rollers, pulleys, and drums. Gearbox oil levels -

Check sight glasses.

Hoist operators pendant connectors secured.

Installed;

Electrical power cables connectors secured.

Condition;

WARNING

Compute with and

CABLE

ACTUATION

OF

CABLE

CUT

CUT

SWITCH ON HOIST CONTROL BOX CAN CUT CABLE, EVEN IF

2-4-A. HOISTINSTALLATION CHECK

If pilot plans to operate hoist with other crewmember in passenger compartment, hoist shall be Installed in forward right position.

OF

SWITCH ON PEDESTALCAN CUT CABLE REGARDLESS OF HOIST PWR SWITCH POSITION.

2-4. INTERIORAND PRESTARTCHECK

NOTE

ACTUATION

CABLE CUT CIRCUIT BREAKER IS

OUT.

CABLE CUT switches (pedestal and hoist) -

Off; covers safetied.

Safety vests, tether straps, hoisting slings, and litters - Condition; secured or stored.


BHT-412-FMS-7

2-4-B. HOIST OPERATION CHECK

HOIST PWR switch - OFF.

Cargo doors and hinged panels - Secured open or removed.

NON ESNTLBUS switch - NORMAL. BATTERY switches - OFF.

HOIST PWR. CONT and CABLE CUT circuit breakers -

In.

NOTE

BATTERY switches external power).

ON (or connect

NON ESNTL BUS switch -

MANUAL.

ICS - Check intercom between pilot and hoist operator using hoist pendant ICS trigger and HOT MIC switch (right ICS box only).

HOIST PWR switch - ON; check that green (power on) light and amber 20 FOOT CAUTION lights on hoist control box illuminate. Hoist OVERTEMPwarning lights to test.

Press

Ground crewmember should be instructed to discharge helicopter static electricity before attaching load to hoist when possible.

2-6. SYSTEMSCHECK

Cargo doors and hinged panels - Secured open or removed. CABLE CUT switches (pedestal and hoist) - Off; covers safetied.

HOIST PWR, CONT, and CABLE CUT circuit breakers - In.

2-6-A. BEFORE TAKEOFF

CAUTION

MAINTAIN TENSION ON HOIST CABLE WHILE REELING IN AND OUT TO PREVENTSLACK.

HOIST and BOOM switches (pilot and operator) - Actuate to check all hoist functions for proper operation. Check that pilot HOIST switch overrides operator pendant HOIST switch.

Safety vests and straps - On and secured to helicopter. Gloves - On.

STEP switch (if installed) - STOW.

2-9. IN-FLIGHTOPERATIONS

Maximun hoist load shall be determined prior to each hoist operation.

Hoist cable - Check for corrosion, kinks, flat spots, fraying, or broken strands.

NOTE

Up limit switch actuator - Raise while hoist is reeling in and check hoist motor stops; then release and check hoist resumes operation.

The Hight-Velocity Diagram is not a limitation for internal hoist operations under an appropriate operating certificate.

Reduce hoist speed as cable approaches up limit. Check that hoist stops when hook reaches up limit without excess tension on

HOIST PWR switch - ON.

cable.

Hoist - Stowed for flight; hook restraint secured. 31


BHT-412-FMS-7

WARNING

HOIST OPERATOR SHALL BE SECUREDTO HELICOPTERWITH AN APPROVED SAFETY HARNESS DURING HOIST OPERATIONS.

Establish hover over hoist operation area.

FOULING ON FUSELAGE OR LANDING GEAR.

Maintain zero ground speed until load is clear of obstructions.

BOOMswitch - IN to swing hoist boom and load into cabin, If possible. Takeoff into wind, if possible, allowing adequate hoist load clearance over obstacles if load is not internal.

Hoist hook restraint - Removed.

BOOM switch (or pilots HOIST switch) OUT. NOTE

Each hoist operation performed is defined as reeling hoist cable out and then in while hovering with any weight on hoist, regardless of whether the hoist was used for training or an actual rescue.

The pilot must record each operation in the penalty CG region. For each hoist operation performed within penalty CG region, four (4) additional hours of usage must be logged against the main rotor yoke, mast and lower cone seat. HOIST switch - DOWN.

Discharge static electricity when possible, and connect hook to load, observing allowable hoist load. NOTE

As hook nears the up or down

limits, hoist speed automatically slows.

HOIST switch - UP.

CAUTION

USE CARE TO PREVENTCABLE, HOOK, AND LOAD FROM

CAUTION

AIRSPEED WITH EXTERNAL LOAD IS LIMITED BY CONTROLLABILITY. CAUTION SHOULD BE EXERCISED WHEN CARRYING AN EXTERNALLOAD. HANDLING CHARACTERISTICS MAY BE AFFECTEDBY THE SIZE, WEIGHT, AND SHAPE OF LOAD.

Airspeed - As required for adequate controllability, not to exceed limits for hoist operations (20 or 60 KIAS, as applicable).

2-13. LITTER HOISTING

When emergency transportation of a patient by litter is essential, every effort should be made to land the helicopter for litter loading. Litter hoisting can be hazardous and should be accomplished only when a landing is not feasible and the condition of the patient precludes the use of the personnel hoisting sling.

In addition to all other procedures contained herein, the following shall apply to litter hoisting operations.


2-13-A. EMPTY LITTER WARNING

HOISTING OR LOWERING AN EMPTY LITTER IN OPEN POSITION IS PROHIBITED. AN EMPTY LITTER CAN OSCILLATE UNCONTROLLABLY IN ROTOR WASH AND FLY UPWARD, STRIKING FUSELAGE OR TAIL ROTOR.

Prior to hoisting or lowering an empty litter, litter shall be closed and secured with straps. Litter should be suspended in a near-vertical position and sling straps should be drawn tight.

2-13-B. LOADED LITTER

BHT-412-FMS-7

SAFETY STRAPS.

HOIST HOOK CATCH SHALL BE SECURED WITH SAFETY PIN PRIOR TO HOISTING LITTER PATIENT.

Litter sling straps should be adjusted so that litter is 24 to 28 inches (61 to 71 centimeters) below hoist hook. NOTE

If litter is suspended too far below hook, litter cannot be loaded in helicopter with hoist hook at up limit.

CAUTION

A LOADED LITTER CAN ROTATE ABOUT CABLE DURING HOISTING. HOIST OPERATOR MAY HAVE TO GRASP LITTER SLING STRAPS TO CONTROL ROTATION AS LITTER APPROACHES LANDING GEAR.

WARNING

LITTER PATIENT SHALL BE SECURED TO LITTER WITH

33


BHT-412-FMS-7

ON CONTROL BOX ON EACH SIDE OF HOIST

ATTACHED TO HOIST

ONCONTROL BOX ON TOPOF HOIST

Figure 2-1.

34

Internal hoist controls

T N Section3 I R P N O R I O T F C U T O OD N R E P E E R Section4 SCR N G O N I R W O E I F V NLY O BHT-412-FMS-7

FAA APPROVED

3-15. HOIST LOAD JETTISON

To jettison hoist load in an emergency, actuate CABLE CUT switch (located on pedestal or hoist).

In the event of failure of CABLE CUT switch, sever cable with manual cable cutter (stowed in pouch on hoist).

3-15-A. HOISTOVERTEMP WARNINGLIGHT

In the event that the OVERTEMPwarning light (located on hoist control box) illuminates, continue present operation until hoist is reeled in. Leave HOIST PWR switch ON (for cooling fan operation) and allow hoist to cool. When OVERTEMPlight extinguishes, hoisting may be resumed as desired.


35/36

T N I Section 5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

5-11.

WEIGHT EMPTY

CHART

BHT-412-FMS-7

NOTE Allowable hoist loads must be computed when AUX Fuel kits are

installed.

The Weight empty chart for internal hoisting operations

is shown in figure 5-1.

Refer to the maintenance manual for

additional information.

NOTE

Allowable hoist load must be computed when weight empty is not within specified guidelines.

37


MANUFACTURER'S DATA

412 Weight Empty Chart for Internal Hoist Operations

8500

(3855.5)

8000 (3628.7)

6500 (2948.3)

140 (3556)

141 (3581)

142 (3607)

143 (3632)

144 (3658)

FUSELAGESTATION- INCHES(MILLIMETERS)

Figure 5-1.

38

Weight empty chart

145 (3683)

412FS-7--1

T N I R P N O R I O T ROTORCRAFT F C U MANUAL T FLIGHT O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-8

MODELS

SUPPLEMENT FOR

LITTER KIT OPERATIONS (412-706-006)

CERTIFIED

SEPTEMBER29, 1981

This supplementshall be attached to the Models 412 or 412EP Flight Manual when the 412-706-006 Litter Kit has been installed.

The informationcontainedhereinsupplementsthe information of the basic Flight Manual. For Limitations,Procedures,and PerformanceData not containedin thissupplement,consultthe basic FlightManual.

Bell Helicopter

A Subsidiary of Textron Inc NOTICE COPYRIGHT

POST OFFICE BOX 482 * FORT WORTH. TEXAS 76101

COPYRIGHT 1994

BELL HELICOPTER TEXTRON INC. ANDBELLHELICOPTER TEXTRON A DIVISIONOFTEXTRON CANADA LTD ALLRIGHTSRESERVED

REISSUE

-

5 OCTOBER

1994


BHT-412-FMS-8

LOG OF REVISIONS

Original .... Reissue .... Reissue ....

0 0 0

.... 29 Sep 81 .. .08 May 89 .... 05 Oct 94

LOG OF PAGES

FLIGHT MANUAL

MANUFACTURER'SDATA

Revision No.

Page

Title - A ...................

0

1--2......................

0

......................

0

Page

3/4 .......................

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT.WORTH, TEXAS 76193-0170

NOTE: Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

A

Revision No. 0



INTRODUCTION

The Litter Kit provides three litters and the provisions for installing up to three litters in the helicopter. A cabin attendant seat is also included in the kit.

i/ii

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R Section 2 SCR N G O N I R W O E I F V NLY O FAA APPROVED

Section 1 and2

BHT-412-FMS-8


MINIMUM

Actual weight change shall be determined after the litter(s) are installed and ballast readjusted if necessary, to return empty weight CG within allowable limits.

The minimum flight crew for litter operations shall consist of a pilot and a second crewmember or cabin attendant, both of whom shall be trained in and capable of assisting in litter patient emergency evacuation procedures.

LITTER LOADING

Secure patients to litters, then load litters aboard the helicopter in sequence from top to bottom. When only two patients are carried, they should occupy the top and center litter positions. When only one patient is carried, the center litter should be used.

LITTER UNLOADING

NOTE

Normal unloading procedures apply when either passenger door can be

FLIGHT CREW

opened.Refer to Section 3, Emergency Procedures, for unloading procedures when cabin door cannot be opened.

Open cabin door and unload litters and patients from the helicopter in sequence from bottom to top.

Litters to be handled by one person inside cabin and one person outside cabin.

T N I Section 3 R P N O R I O CT F U T O OD N R E P E E R R C S G Section 4ON N I R W O E I F V NLY O Section 3 & 4

FAA APPROVED

BHT-412-FMS-8

UNLOADING

THROUGH

Unstrap

EMERGENCY EXITS

In the event that cabin doors cannot be opened, litter patients shall be unloaded through emergency pop-out windows. After all litter patients have been removed, ambulatory patients may then exit.

litter

and remove

window

by

Raise top and center litter ends near open window and engage center litter in brackets for top litter. Disconnect bottom litter. Raise bottom litter at end near open window and rest handles on the lower surface of the window opening. Unstrap patient and slide patient up litter and through window opening.


2

on center

Disconnect top litter at end near open window and lower end to rest on center litter. Removepatient retention straps and slide patient down litter and out through window opening.

NOTE

Remove emergency pop-out pushing at corners as marked.

patient

patient through window opening.

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section 1

MANUFACTURER'SDATA

BHT-412-FMS-8

MANUFACTURER'SDATA

TABLE OF MOMENTS

TABLE OF MOMENTS

(IN-LB)

(kg * mm) 100

Weight (Pounds)

LITTERPATIENT Loaded Laterally F.S. 117

100 110 120 130 140 150 160 170 180 190 200 210 220

11700 12870 14040 15210 16380 17550 18720 19890 21060 22230 23400 24570 25740

Weight (K.G.)

LITTERPATIENT Loaded Laterally 2972 mm

50 55 60 65 70 75 77.1 80 85 90 95 100 105 110

1486.0 1634.6 1783.2 1931.8 2080.4 2229.0 2291.4 2377.6 2526.2 2674.8 2823.4 2972.0 3120.6 3269.2

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-9.3 AND 9.4

MODELS


SUPPLEMENTFOR EXTERNALCARGOOPERATION (212-706-103) 36020 -

36086

AND

36087 AND SUB

CERTIFIED 25 SEPTEMBER1991

This supplement shall be attached to the Model 412 Flight Mnaual (BHT-412-FM-3) or Model 412EP Flight Manual (BHT-412-FM-4) when the 212-706-103 External Cargo Suspension Hook has been installed.


Bell Helicopter INC. BELLe HELICOPTER ANDBELLHELICOPTER TEXTRON INC. ADIVISIONOFTEXTRONCANADALTD. ALLRIGHTS RESERVED

REISSUE -

15 SEPTEMBER 1995

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-9.3AND 9.4

NOTICEPAGE





FAA APPROVED

LOGOF REVISIONS

..........

0..............

25 Sep 91

Reissue...........0

...............

05 Oct 94

Reissue ...........

0 ...............

08 Oct 93

Reissue ...........

0..............

15 Sep 95

Original

LOGOF PAGES

REVISION NO.

PAGE

MANUFACTURER'S DATA

FLIGHT MANUAL

Title - NP .............................. A - B .................................. i/ii ....................................... 1--10 ..................................

REVISION NO.

PAGE

0 0 0 0

11 -12.................................0 13/14 ...................................

0

NOTE

Revised text Is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. A

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W 0 O E I F V NLY O BHT-412-FMS-9.3 AND 9.4

FAA APPROVED


Original ........... Reissue ...........

0 .............. 0 ...............

25 Sep 91 08 Oct 93

APPROVED:

MANAGER


Reissue...........0 Reissue ...........

............... ..............

05 Oct 94 15 Sep 95


FAA APPROVED

GENERAL INFORMATION INTRODUCTION

The External Cargo Suspension Hook, when installed, will permit the operator to utilize the helicopter for transportation of external cargo, when operated by a qualified pilot.

i/ii

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-9.3 AND 9.4

FAA APPROVED


Operation of the helicopter with no load on the external cargo suspension hook is under the standard authorized 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 the operators discretion.

The rear view mirror shall be covered or removed for night flight.

1-6. WEIGHT AND CENTER OF GRAVITY

Actual weight change shall be determined after cargo hook is Installed and ballast readjusted, if necessary, to retain empty weight CG within allowable limits.

1-7.

AIRSPEED

VNE is 80 KIAS at or below 10,000 feet density altitude for all gross weights with external cargo on suspension unit. Above 10,000 feet decrease VNE 2.5 knots per 1000 feet.

1-3-C. VFROPERATION

With a load attached to the suspension assembly, operation shall be conducted in accordance with appropriate operating rules for external loads under VFR conditions.

1-3-D. IFR OPERATION

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

CAUTION

THE AIRSPEED WITH EXTERNAL BY IS LIMITED CARGO CONTROLLABILITY. CAUTION SHOULD BE EXERCISEDWHEN CARRYING EXTERNAL CARGO, THE HANGING AS CHARACTERISTICS MAY BE BY THE SIZE, AFFECTED WEIGHT, AND SHAPE OF THE CARGO LOAD.

Light weight, high drag loads require a swivel connector between the cargo hook and the sling to prevent unstable oscillations in flight above 20 KIAS.

1


FAA APPROVED

PLACARDS AND MARKINGS

OCCUPANCY LIMITED TO CREW WITH EXTERNAL LOAD CLASS B LOADING APPROVED

(Located on forward right side of overhead console)

DO NOT OPERATE HEATER ABOVE 21 DEG C OUT AIR TEMP FLOATS

CARGORELEASE

OPERATIONAL LIMITS

(Located on upper center part of instrument

panel)

LIMIT EXTERNAL LOAD 4500 LBS SEE FLIGHT MANUAL FOR LOADING INSTRUCTIONS AND FLIGHT LIMITATIONS

(Located on under side of helicopter near suspension assembly)

412-FMS-9.3-1-1

Figure 1-1.

Placards and markings


FAA APPROVED

Section2

2-2. FLIGHTPLANNING

Instruct ground personnel to discharge before helicopter static electricity attaching cargo by touching the airframe with a ground wire, or if a metal sling is used, the hookup ring can be struct against the cargo hook. If contact has been lost after initial grounding, the should be electrically helicopter regrounded and, if possible, contact maintained until hookup is complete.

WARNING

USE OF INAPPROPRIATELY SIZED LOAD RINGS MAY RESULT IN LOAD HANG-UP WHEN LOAD RING IS TOO SMALL OR INADVERTENTLOAD RELEASE IF LOAD RING IS TOO LARGE.

Check that only one primary ring is captured in load beam and only one secondary ring with correct cross-section dimension is captured in primary ring.

I nstruct ground personnel to check primary and secondary load rings for condition and proper size (Table 2-1). Check for proper rigging and configuration (Figure 2-1).

Table 2-1. Ring Size - Cargo Hook P/N SP1709-62

PRIMARY RING

INSIDE DIAMETER

MAXIMUM CROSSSECTION OF SECONDARY MEMBER

PRIMARY RING

CROSS-SECTION

3.0 to 3.1 in. (76.2 to 78.74 mm)

0.625 in.

1.0 in.

(15.9 mm)

(25.4 mm)

3.1 to 4.0 In. (78.74 to 101.6 mm)

0.750 in. (19.0 mm)

1.0 in.

(25.4 mm)

3


BHT-412-FMS-9.3 AND 9.4

CORRECTRIGGING

*

PRIMARYRING CROSS-SECTION

* MAXIMUM OF CROSS-SECTION MEMBER SECONDARY

*

DIMENSIONS ARE SHOWN IN TABLE 2-1

TO LOAD(S)

INCORRECT RIGGING

INCORRECT RIGGING

412FMS-9.3-2-1

Figure 2-1.

Effective loading practices


FAA APPROVED

2-8. TAKEOFF

2-3. PREFLIGHTCHECK

Cargo suspension assembly - Condition and security and warning plate is visible. Rear view mirror (if installed) and clean.

Secure


NOTE

Better directional control may be realized by avoiding relative winds from the right front while performing quadrant external cargo operations.

Hover helicopter at sufficient height to allow ground personnel to discharge static electricity and to attach cargo sling to cargo hook.

CARGOHOOK REL circuit breaker - In. Battery BUS 1 switch -

ON.

CARGO REL switch (overhead) - ARM; check CARGO RELEASE ARMED caution light illuminates. Cyclic CARGO RELEASE button Depress and hold; pull down on cargo hook; hook should open. Release button and cargo hook; hook should close and lock.

Cargo release pedal (between directional control pedal(s) - PUSH and hold; pull down on cargo hook; hook should open. Release pedal and cargo hook; hook should close and lock. NOTE

The pedal release will function regardless of CARGO REL switch position.

CARGO REL switch -

NOTE

Attachment of cargo sling to the hook can be observed by means of the rear view mirror.

Ascend vertically directly over cargo, then slowly lift cargo from surface. Pedals control.

Check for adequate directional

Hover power - Check torque required to hover with external load. NOTE

The Height - Velocity Diagram is not a limitation for external cargo operations under an appropriate operating certificate.

Takeoff into the wind If possible, allowing adequate sling load clearance over obstacles.

OFF.

Battery BUS 1 switch - OFF.

2-7. BEFORE TAKEOFF

CARGOREL switch - ARM; check CARGO light ARMED caution RELEASE Illuminates. 5

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O Section3 N I R W O E I F V NLY O BHT-412-FMS-9.3AND 9.4

FAA APPROVED

2-9. IN-FLIGHTOPERATIONS NOTE

CARGO REL switch (overhead) prior to final approach.

Control movements should be made smoothly and kept to a minimum to prevent oscillation of sling load.

CARGO REL switch (overhead) desired.

2-10. DESCENTAND LANDING

-

ARM

Flight path and approach angle - As required for wind direction and obstacle clearance.

As

NOTE

The pedal release will function regardless of CARGO REL switch position.

Airspeed - Within limits for adequate controllability of rotorcraft and load configuration.

Execute approach to a hover with cargo clear of the surface. When stabilized at a hover, descend slowly until cargo contacts surface. Maintain tension on sling. Cyclic CARGO RELEASE button Depress to release sling from hook.

-

NOTE

Release of sling load from the hook can be confirmed visually through rear view mirror.

Flight path - As required to avoid flight with external load over any person, vehicle or structure.

3-15. CARGO FAILS TO RELEASE ELECTRICALLY

Maintain tension on sling.

Cargo release pedal (between directional control pedals) -

In the event that cargo hook will not release the sling when the CARGO RELEASE button Is depressed, proceed as follows:

PUSH.

T Section4 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-9.3AND 9.4

FAA APPROVED

4-1. INTRODUCTION

No change from basic Flight Manual performance with no load attached to cargo hook. NOTE

operations, refer to figure 4-1, critical relative wind azimuths.

In ground effect hover ceiling chart for external cargo operations is presented in figure 4-2.

Refer to basic Flight Manual for out of ground effect hover ceiling charts during external cargo operations.

Performance may be affected by the size and shape of the external load.

4-5. HOVERCEILING NOTE

When using any hover ceiling cargo chart for external

7


270

FAA APPROVED

°

412-FMS-9.3-4-1

Figure 4-1. Critical relative wind azimuths


FAA APPROVED

HOVER CEILING

IN GROUNDEFFECT

SKID HEIGHT 4 FEET HEATERON OR OFF

TAKEOFF POWER ENGINERPM 100%

° -40 TO 52°C

150 AMPSEACH GENERATOR

NOTE: THIS DATA VALID FOR ZERO WIND OR WIND OUT SIDE OF THE CRITICAL WIND AZIMUTH AREAS OAT-°C

14.000 FT. DEN. ALT. LIMIT

-10 -20

WEIGHT GROSS GROSS WEIGHT

Figure 4-2. Hover ceiling in ground effect (Sheet 1 of 2)

412.FMS-9.3-4-29 412-FMS-9.3'4-2-1

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-9.3 AND 9.4

FAA APPROVED

HOVER CEILING

IN GROUND EFFECT

MAXIMUMCONTINUOUS POWER ENGINERPM100% GENERATOR 150 AMPSEACH

SKIDHEIGHT4 FEET HEATER ONOROFF -40 ° TO52°C

NOTE: THISDATAVALIDFORZEROWINDORWINDOUTSIDEOFTHECRITICALWINDAZIMUTHAREAS. 14,000 FT DEN.ALT. LIMIT

-30 -40

-20

MAX OAT

TORQUE

LIMIT

OAT

°C

3.5

4.0 4.5 5.0 GROSS WEIGHT

5.4 kg X 1000

412-FMS-9.3-4-2-2

Figure 4-2.

10

Hover ceiling In ground effect (Sheet 2)

T Section 5 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

BHT-412-FMS-9.3 AND 9.4

The External Cargo Loading Tables (tables 5-1 and 5-1M) present moments for external loads suspended from the cargo hook at fuselage station 138.0 (3505 mm).

11


Table 5-1.

MANUFACTURER'S DATA

External cargo loading table (English)

EXTERNAL CARGO LOADING TABLE -

Cargo Weight (Lb)

Moment F.S. 138.0

ENGLISH

Cargo Weight (Lb)

Moment F.S. 138.0

Cargo Weight (Lb)

Moment F.S. 138.0

50 100 150 200 250

6900 13800 20700 27600 34500

1800 1850 1900 1950 2000

248400 255300 262200 269100 276000

3550 3600 3650 3700 3750

489900 496800 503700 510600 517500

300 350 400 450 500

41400 48300 55200 62100 69000

2050 2100 2150 2200 2250

282900 289800 296700 303600 310500

3800 3850 3900 3950 4000

524400 531300 538200 545100 552000

550 600 650 700 750

75900 82800 89700 96600 103500

2300 2350 2400 2450 2500

317400 324300 331200 338100 345000

4050 4100 4150 4200 4250

558900 565800 572700 579600 586500

800 850 900 950 1000

110400 117300 124200 131100 138000

2550 2600 2650 2700 2750

351900 358800 365700 372600 379500

4300 4350 4400 4450 4500

593400 600300 607200 614100 621000

1050 1100 1150 1200 1250

144900 151800 158700 165600 172500

2800 2850 2900 2950 3000

386400 393300 400200 407100 414000

4550 4600 4650 4700 4750

627900 634800 641700 648600 655500

1300 1350 1400 1450 1500

179400 186300 193200 200100 207000

3050 3100 3150 3200 3250

420900 427800 434700 441600 448500

4800 4850 4900 4950 5000

662400 669300 676200 683100 690000

1550 1600 1650 1700 1750

213900 220800 227700 234600 241500

3300 3350 3400 3450 3500

455400 462300 469200 476100 483000

(TABLE

I.D. 911494)


MANUFACTURER'S DATA

Table 5-1M.

External cargo loading table (Metric)

EXTERNAL CARGO LOADING TABLE -

Cargo Weight (kg)

Moment 3505 mm (kg - mm) 100

METRIC

Cargo Weight (kg)

Moment 3505 mm (kg - mm) 100

40 80 120 160 200

1402 2804 4206 5608 7010

1240 1280 1320 1360 1400

43462 44864 46266 47668 49070

240 280 320 360 400

8412 9814 11216 12618 14020

1440 1480 1520 1560 1600

50472 51874 53276 54678 56080

440 480 520 560 600

15422 16824 18226 19628 21030

1640 1680 1720 1760 1800

57482 58884 60286 61688 63090

640 680 720 760 800

22432 23834 25236 26638 28040

1840 1880 1920 1960 2000

64492 65894 67296 68698 70100

840 880 920 960 1000

29442 30844 32246 33648 35050

2040 2080 2120 2160 2200

71502 72904 74306 75708 77110

1040 1080 1120 1160 1200

36452 37854 39256 40658 42060

2240 2268

78512 79493

(TABLE

I.D. 911495)

13/14

BHT-412-FMS-10.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O (BHT-412-FM-3)when operating in Category A conditions and PT6T-3BEor PT6T-BGengines are installed.

The information contained herein supplementsinformation of the basic Flight Manual.For limitations, Procedures,and Performance Data not contained in this supplement, consult the basic Flight Manualor other applicablesupplements.

COPYRIGHT NOTICE

Bell Helicopter

COPYRIGHT 1999 BELL HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON INC. A DIVISIONOF TEXTRON CANADA LTD. ALL RIGHTS RESERVED

POST OFFICE BOX 482

A Subsidairy of Textron Inc. FORT WORTH, TEXAS 78101

REISSUED -

23 APRIL 1999

REVISION 1 - 24 NOVEMBER 1999


NOTICEPAGE



LOG OF REVISIONS

Original .. ..................... 0......................5 OCT 94 Reissue ....................... 0................... 23 APR 99 Revision ...................... 1.................... 24 NOV 99

LOG OF PAGES

REVISION NO.

PAGE

PAGE

REVISION NO.

FLIGHT MANUAL

Title............................................................ 1 NP .............................................................. 0 A - B ........................................................ 1 i- ii .......................................................... 0 1-1 .............................................................. 1 1-2 - 1-4 ...................................................0 2-1 - 2-2 ...................................................0 2-3 /2-4......................................................0 3-1 ..............................................................1 3-2 - 3-6 ................................................... 0 3-7 /3-8 ...................................................... 0 4-1 - 4-2 ................................................... 1 4-3 - 4-9 ...................................................0 4-10 - 4-18 ............................................... 1 4-19/ 4-20.................................... 1

NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. Rev. 1

A


BHT-412-FMS-10.3


Original ........................ 0 ..................... 5 OCT 94 Reissue ....................... 0.................. 23 APR 99 Revision ....................... 1 ................... 24 NOV 99

APPROVED:


Rev. 1

DATE:

NOV2 4 1999


INTRODUCTION

The information contained in this supplement is for category "A" operations. For limitations, normal procedures, emergency and malfunction procedures, and performance data not contained in this supplement, consult the appropriate sections of this flight manual.

GENERALINFORMATION

DEFINITIONS:

CATEGORY"A" TAKEOFF

-

Operation of the helicopter in such a manner that if one engine fails at any time after the start of the takeoff, the helicopter can: 1. At or prior to CDP, return to and safely stop on the takeoff area; or 2. At or after CDP, climb out from point of failure and attain single engine forward flight.

CATEGORY"A" LANDING

-

Operation of the helicopter in such a manner that if one engine fails at any time after the start of a landing approach the helicopter can:

1. At or after LDP, continue the approach and safely land and stop on the clear heliport; or 2. At or prior to LDP, climb out from point of failure and attain single engine forward flight.

CRITICALDECISIONPOINT

-

The last point in the takeoff path at which a rejected takeoff can be assured, and the first point at which a completed takeoff can be assured.

LANDINGDECISIONPOINT

-

That point on the landing profile after which the helicopter is committed to landing.


COMPLETEDTAKEOFF DISTANCEREQUIRED

-

The horizontal distance from the start of the prescribed takeoff procedure to a point at least 35 feet above the takeoff surface where VToSSand a positive rate of climb are attained following an engine failure occurring at or after CDP.

REJECTEDTAKEOFF DISTANCE REQUIRED

-

The horizontal distance from the start of the prescribed takeoff procedure to the point where the helicopter is brought to a safe stop on the designated surface following an engine failure occurring at or prior to CDP.

TAKEOFF FLIGHT PATH

-

The distance traveled from where the aircraft at or above 35 feet AGL to 1000 feet reaches VTOSS AGL.

TAKEOFF SAFETYSPEED

-

The airspeed that will assure the required climb performance with one engine inoperative.

LANDING DISTANCE REQUIRED

-

The horizontal distance necessary to achieve a takeoff flight path at VTOssand an altitude of 35 feet or higher, with one engine inoperative at or prior to LDP; or the horizontal distance necessary to land the helicopter without further incident, with one engine inoperative at or after LDP.

BALKED LANDING

-

The discontinuation of a landing approach and the initiation of a climbout. Category "A" balked landing capability following an engine failure is assured at or prior to LDP.

ABBREVIATIONS:

Ground Level

AGL

-Above

CDP

-Critical

Decision Point

CDT

-Critical

Decision Time

GROC

-Gross

LDP

-Landing

ii

Rate of Climb

Decision Point

VMINIFR -Minimum

Airspeed for IFR

Safety Speed

VToSS (V2)

-Takeoff

V,

-—Best Rate of Climb Speed

WAT

- Weight-Altitude-Temperature

T * Section 1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-10.3


Category A operationis approvedfor day/night VMC, non-icing conditions. Night takeoff and landing may be accomplished with adequate lighting.

1-5. CONFIGURATION

Standardlanding gearor high skid gearwith or without emergencyfloats (floats stowed).

Pratt and Whitney PT6T-3BE or PT6T-3BG engines shall be installed.

1-6. WEIGHT AND CENTEROF GRAVITY

1-6-C. TAKEOFFAND LANDING WEIGHTVS ALTITUDE Refer to Gross Weight-Altitude-AmbientAir TemperatureLimits Charts (figure 1-1).Charts designated Part A may be used for gross

weights to 10,000 pounds (4636 kg). Part B charts maybe used for gross weights to 10,800 pounds (4899kg). Part C charts may be used for grossweights to 11,900pounds (5398kg). Interpolation of data between charts for different parts is not permitted.Testinghas not been conducted in areas betweenParts A, B, and C.

1-7. AIRSPEED

1-7-A. CROSSWIND LIMITATIONS

The crosswind limit for takeoff and landing is 20 knots. Refer to the unfactored Headwind PERFORMANCE Component Chart in subsection. Takeoff or landing downwind or with quartering tailwinds is prohibited.

1-8. ALTITUDE

Maximumaltitude for takeoff and landing is 4000feet pressure altitude.

Rev.1

1-1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O APPROVED

WEIGHT - ALTITUDE

- TEMPERATURE PART A

FOR TAKEOFF AND LANDING GW TO 10,000

VTOSS = 40 KIAS

LBS (4636 kg)

4000

3000-

GROSS WEIGHT

Figure1-1. Grossweight-altitude-ambient air temperature limits chart-takeoff and landing (Sheet 1 of 3) 1-2


FAA APPROVED

WEIGHT - ALTITUDE

VTOSS = 46 KIAS

- TEMPERATURE FOR TAKEOFF AND LANDING PART B

GW TO 10,800 LBS (4899 kg

4000

Figure 1-1.

Grossweight-altitude-ambient air temperature limits chart-takeoff and landing (Sheet 2 of 3) 1-3


FAA APPROVED

WEIGHT - ALTITUDE - TEMPERATURE FOR TAKEOFF AND LANDING PART C

VTOSS = 55 KIAS

4000

3000

2000-

Figure 1-1.

Grossweight-altitude-ambient air temperature limits chart-takeoff and landing (Sheet 3 of 3)

T N * Section 2 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-10.3

FAA APPROVED

NOTE

CDP height is determined by reference to the pilots barometric altitude. Indicated altitude with collective full down on the takeoff surface is used as a ground level reference.

After attaining CDP, accelerate the helicopter to 65 KIAS and continue the climb.

2-10. DESCENTAND LANDING 2-10-A. STANDARDTYPE LANDING NOTE

skid height and note torque.

AdjustADI pitch bar to indicate level.

Initiate a takeoff from hover using a MAST TORQUE of 10% above that required to hoverand ten degreesnose down altitude. NOTE

Do not exceed MASTTORQUE,ITT, or GASPRODRPM(N1)limits. Maintain pitch altitude as the helicopter movesforwardto achievethe correct Critical Decision Point (CDP)shown on the takeoff flight path profile diagram(figure 2-1).

A standard type landing is initiated from a Landing Decision Point (LDP) of 40 KIAS andan altitude of 100feet (30.5 meters) above the runway, either in level flight or with a rate of descentof not morethan 500feet per minute (figure 2-2).

Flight controls-Adjust friction to desired level. GOVswitches-AUTO.

Throttles-Fully open.

ENG-100% RPM(N2).

FORCETRIMswitch-As desired.

STEPswitch-As desired.

Altimeter-Set to nearestreporting station. SEAT BELT and NO SMOKE sign-As desired.

2-1

T N I r R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-10.3

TAKE-OFF FLIGHT PATH PROFILE

2 1/2 MINUTE POWERRATING AFTER POWERFAILURE. THEN REDUCE TO 30 MINUTE POWERAT 500 FT (152 m) AGL.

HEIGHT

DISTANCE

2-2

CLIMB TO 1000 FT (305m) AT VTOSSMINIMUM


FAA APPROVED

LANDING FLIGHT PATH PROFILE

35 FT. -(107MI HEIGHT

Figure 2-2.

Landing flight path profile

2-3/2-4

T N · Section3 I R I P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-10.3

FAA APPROVED

* DURING TAKEOFFAFTER CRITICAL

3-1. INTRODUCTION

DECISION POINT (CDP)

Tables3-1through 3-4list panelwording, fault conditions, and corrective actions for emergencies and malfunctions that might occur during takeoff after CDP,during landing prior to LDP,andduring landing after LDP.

3-3. ENGINE

3-3-A. SINGLEENGINE FAILURE

CAUTION

DURING TAKEOFF PRIOR TO CRITICAL DECISION POINT (CDP)

An engine failure prior to reaching CDP will necessitate a landing back to the takeoff surface. If height permits, a positive deceleration to reduce forward airspeed is required. As the helicopter descends, it should be leveled and the collective should be used as required to cushion the landing. Some forward ground speed Is normally required at touchdown. Maintain control of the helicopter.

Collective - Adjust to maintain ROTOR RPMand OEI power limits. Flare to reduce ground speed.

Assume landing touchdown.

attitude

In the event of an engine failure following CDP, airspeed should be increased to the takeoff safety speed (VToss)or maintained, whichever Is higher. Climb out to 500 feet (152 meters) above the takeoff surface and accelerate to 65 KIAS. Reduce power to 30 minute limit (PT6T-3BEengines). I

before

DURING COLD WEATHER CAREFULLY OPERATIONS, MONITOR TORQUE OF THE NORMAL ENGINE WHEN ONE ENGINEFAILS OR IS SHUT DOWN IN FLIGHT. NOTE

During takeoff, after CDP, it is permissible to droop ROTOR RPM to 91% during the transition from twin engine to single engine flight an engine failure. following ROTOR RPM should be regained to normal operating range at or before attaining appropriate best rate of climb speed.

Maintain control of the helicopter.

Collective - Adjust to maintain ROTOR RPM and OEI power limits.

Throttle (affected engine) - Closed.

Airspeed - If below VTOSS, smoothly increase to VTOSsand initiate a climb.

Complete shutdown of affected engine.


Close.

Rev.1

3-1


FAA APPROVED

Complete shutdown of affected engine. ENG (unaffected engine) RPM (N2).

Set to 100%

is accomplished using up to the maximum power of the remaining engine while maintaining rotor speed within limits. Maintain control of the helicopter.

DURING LANDING PRIOR TO LANDING DECISION POINT (LDP)

Execute the same procedures as for single engine failure on takeoff after CDP or proceed to LDP and use the procedure below.

DURING LANDING AFTER THE LANDING DECISION POINT (LDP)

Collective - Adjust to maintain ROTOR RPM and OEI power limits. Flare to reduce speed.


attitude


before

Closed.


The helicopter, with an emergency, Is committed to land after LDP. The landing Table3-1.

PANEL WORDING

Warning lights-Takeoff prior to CDP

FAULT CONDITION

CORRECTIVE ACTION

MASTER CAUTIION

Warning or caution light(s) illuminated.

Land immediately.

FIRE PULL (1 or 2)

Fire indication in No. 1 or Land immediately. No. 2 engine compartment. Pull affected FIRE PULL handle. Select MAIN fire extinguisher; if necessary, select RESERVE fire extinguisher.

BAGGAGE FIRE

Smoke in compartment.

ENG OUT (1 or 2)

GAS PRODabnormally low, Land immediately. below 53 ± 2% RPM (N1), on Refer to ENGINEOUT procedures. No. 1 or No. 2 engine.

XMSN PRESSURE

C BOX TEMP

3-2

Land immediately. Inspect tailboom area for damage.

O I L Transmission oil pressure Land immediately. below limit.

XMSNOIL TEMIP C BOX PRESSURE

baggage


O I L Combining gearbox oil pressure below normal. Combining gearbox temperature above limit.

Land immediately. Land immediately. Land immediately.


FAAAPPROVED

Table 3-1.

PANEL WORDING

BATTERY TEMP

Warning lights-Takeoff prior to CDP(Cont)

FAULT CONDITION

CORRECTIVE ACTION

Battery case temperature above limits.

Land Immediately. BATTERY BUS 1 and BUS 2 switches - OFF. WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATIONOF BATTERY TEMP LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIORTO RETURNTO SERVICE.

ROTOR BRAKE


not

Land immediately.

Table 3-2. Warning lights-Takeoff after CDP

PANEL WORDING

FAULT CONDITION

CORRECTIVE ACTION

MASTER CAUTION

Warning or Caution light(s) Illuminated.

Accelerate to VToss. Reset MASTER CAUTION light; take appropriate corrective action as required by Illuminated segment(s).

FIRE PULL (1 or 2)

Fire Indication in No. 1 or Accelerate to VTOSS.Pull affected No. 2 engine compartment. FIRE PULL handle. Select MAIN fire Close throttle of extinquisher. affected engine. Select RESERVE fire extinguisher if necessary. Land as soon as possible.

BAGGAGE FIRE

In Smoke compartment.

ENG OUT (1 or 2)

to VTOSS. Secure GAS PROD abnormally low, Accelerate below 53 ± 2% RPM (N1), on appropriate engine. Refer to ENGINE OUT procedure. Land as soon as No. 1 or No. 2 engine. possible.

baggage


3-3


BHT-412-FMS-10.3

Table3-2.

PANEL WORDING

X MS N PRESSURE

Warning lights-Takeoff after CDP (Cont)

FAULT CONDITION O I L Transmission oil pressure below limit.

XMSNOILTEMP

oil Transmission temperature above limit.

CORRECTIVE ACTION Accelerate to VToss.Reduce power; verify fault on XMSN OIL pressure gage. Land as soon as possible. Accelerate to VTOSS.Reduce power;

on XMSN OIL fault verify temperature gage. Land as soon as possible.


Accelerate to VTOSS'Reduce power;

C BOX TEMP


Accelerate to VTOSS.Reduce power;

BATTERYTEMP


Accelerate

B O X C PRESSURE

verify fault on GEAR BOX pressure gage. Land Immediately. on GEAR BOX fault verify temperature gage. Land as soon as possible. to VTOSs.BATTERY BUS

1 and BUS 2 switches as soon as practical.

OFF. Land

I WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATION OF BATTERY TEMP LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIORTO RETURNTO SERVICE.

ROTOR BRAKE

3-4


not

Accelerate

to VTOSS. Check rotor

brake handle fully up In detent. If light remains on, land as soon as possible.


FAA APPROVED

Table3-3.

PANEL WORDING

Warning lights-Landing prior to LDP

FAULT CONDITION

CORRECTIVE ACTION

MASTER CAUTION

Warning or caution light(s) Reset MASTER CAUTION light; take illuminated. appropriate corrective action as required by illuminated segment.

FIRE PULL (1 or 2)

Fire indication in No. 1 or Pull affected FIRE PULL handle. No. 2 engine compartment. Select MAIN fire extinguisher. Close throttle of affected engine. Select if RESERVE fire extinguisher, Land as soon as necessary. possible.

BAGGAGE FIRE E

Smoke

ENG OUT (1 or 2)

GAS PRODabnormally low, Maintain VTOSS.Secure appropriate below 53 ± 2% RPM(N1), on engine. Land as soon as possible. Refer to ENGINEOUT procedure. No. 1 or No. 2 engine.

XMSN PRESSURE

baggage

O I L Transmission oil pressure below limit.

XMSN OIL TEMP

BOX C PRESSURE

in

compartment.



C BOX OIL TEMIP



Reduce power. Verify fault on XMSN Land gage. OIL pressure immediately. Reduce power. Verify fault on XMSN OIL temperature gage. Land as soon as possible. Reduce power. Verify fault on GEAR BOX pressure gage. Land as soon as possible. Reduce power. Verify fault on GEAR BOX temperature gage. Land as soon as possible.

3-5


FAA APPROVED

Table 3-3. Warning lights-Landing prior to LDP (Cont)

PANEL WORDING

BATTERYTEMP

FAULT CONDITION

CORRECTIVE ACTION


BATTERY switches practical.

BUS 1 and BUS 2 OFF. Land as soon as

WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATIONOF BATTERY TEMP LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIORTO RETURNTO SERVICE.

ROTOR BRAKE


Table 3-4.

PANEL WORDING

not

Check rotor brake handle fully up in detent. If light remains on, land as soon as possible.

Warning lights-Landing after LDP

FAULT CONDITION

CORRECTIVE ACTION

MASTERCAUTION


Land immediately.

FIRE PULL (1 or 2)

Fire indication In No. 1 or Land immediately. Pull affected No. 2 engine compartment. FIRE PULL handle. Close throttle of affected engine. Select MAIN fire extinguisher; If necessary, select RESERVEfire extinguisher.

BAGGAGEFIRE

Smoke In compartment.

ENG OUT (1 or 2)

GAS PROD abnormally low, Land Immediately. below 53 ± 2% RPM (N1), on Refer to ENGINEOUT procedures. No. 1 or No. 2 engine.

XMSN PRESSURE 3-6

baggage

01 L Transmission oil pressure below limit.


Land Immediately.


FAA APPROVED

Table 3-4. Warning lights-Landing after LDP (Cont)

PANEL WORDING

FAULT CONDITION

CORRECTIVE ACTION

XMSN OIL TEMP


Land immediately.

L Combining gearbox oil pressure below normal.

Land immediately.

C BOX TEMP

gearbox Combining temperature above limit.

Land immediately.

BATTERYTEMP


Land immediately. BATTERY BUS 1 and

BOX C PRESSURE

0I

switches -

BUS 2

OFF.

WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATION OF BATTERY

TEMP

LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURNTO SERVICE.

ROTORBRAKE


not

Land immediately.

3-7/3-8


4-1. INTRODUCTION

BHT-412-FMS-10.3

4-5. HOVERCEILING 4-5-A. HEADWIND

The power performancedata presentedin this section is based on engine manufacturers COMPONENT minimum specificationpowerfor the PT6T-3BE The Unfactored Headwind Component chart engineswith installationlosses. and PT6T-3BG (figure 4-1) is provided with an example to determine critical crosswind and corrected headwind for Category "A" takeoff and landings. The headwind component, as The takeoff and landing data presentedin this calculated from the headwind component section is based on tests performedon a level chart, is applied to parts A, B, and C of the asphalt runway 75 feet wide. The minimum TakeoffSpaceRequired charts. runway length for standardtakeoffand landing procedures varies with wind, gross weight, 4-6. TAKEOFFDISTANCE pressurealtitude, and temperature.

4-2. POWERASSURANCE CHECKS

Refer to Section 4 of applicable Basic Flight Manual for power assurance charts to determine if the engine (power sections) can produceinstalledspecification power.

The hover check is performed prior to takeoff. The in-flight check is provided for in-flight monitoring of engine performance. If either engine (power section) does not meet the requirements of the hover or in-flight power assurance check, category "A" performance will not be achievable. The cause of engine power loss, or excessive interturbine temperature(ITT)or GASPRODRPM(N1)shall be determinedas soon as practical. Refer to appropriate engine maintenancemanual.

4-6-A. REJECTED TAKEOFF DISTANCEREQUIRED

The rejected takeoff distance required Is the space necessary to takeoff, climb to CDP, encounter an engine failure at CDP, return to takeoff surface, and stop safely. The rejected takeoff distance required is obtained from either part A, B, or C of the Rejected Takeoff Distance Required charts (figure 4-2).

4-6-B. TAKEOFFSPACE REQUIRED

The takeoff space required Is the horizontal distance required to takeoff, climb to CDP, encounter an engine failure, accelerate toVToss, and climb to 35 feet (10.7 meters) AGL

above the takeoff space. Takeoff space requiredis obtained from either part A, B, or C of the TakeoffSpace Required charts (figure 4-3) using the headwind component from the Unfactored Headwind Component chart (figure 4-1). Rev.1


4-6-C. TAKEOFFFLIGHT PATH

The takeoff flight path begins at the end of Takeoff Space Required, at 35 feet (10.7 meters)AGLor higher,abovethe takeoffspace and at VTOSS.Parts A, B, and C of the Takeoff

Flight Path charts (figure 4-4 for PT6T-3BE engines and figure 4-4A is for PT6T-3BG engines) providedata for 35 to 500feet (10.7to 152 meters) and 500 to 1,000feet (152to 305 meters) AGL. These charts provide altitude gain for each 100feet (30.5 meters)horizontal distance traveled.

4-9. LANDING

4-9-A. LANDING SPACE REQUIRED

Landing space required is the distance necessary to come to a stop from LDP with one engine inoperative.

4-2

Rev.1

FAA APPROVED

Landing space required is obtained from the Landing Space Required chart (figure 4-5).

4-9-B. CORRECTEDLANDING DISTANCE

Corrected landing distance from LDP is landing distance corrected for wind factor.

The headwind component is obtained from calculation of the Unfactored Headwind Componentchart (figure 4-1), and applied to the Landing SpaceRequiredchart (figure 4-5) to obtain corrected landingdistance.


FAAAPPROVED

0

Z

40

EXAMPLE ° 1. TAKE OFF HEADING ...................................................... 170 ° 2. REPORTED WIND DIRECTION ........................................... 200 ° 3. WIND DIRECTION, DEGREES FROM TAKE-OFF HEADING ................. 30 4. REPORTED WIND SPEED ..................................... 20 KNOTS 5. ENTER CHART AT REPORTED WIND SPEED. POINT A 6. FOLLOW THE SHAPE OF THE CURVED LINES, TO WIND DIRECTION. DEGREES FROM TAKE-OFF HEADING. POINT B. 7. PROCEED HORIZONTALLY TO THE HEADWIND COMPONENT SCALE AND READ HEADWIND COMPONENT ...... 17 KNOTS 8. TAILWINDS HAVE NOT BEEN DEMONSTRATED

Figure 4-1.

Unfactored headwind component chart 4-3


FAA APPROVED

REJECTED

TAKE-OFF DISTANCE

REQUIRED

PART A

GW 7.500 TO 10,000 LB (3402 TO 4536 Kg)

VTOSS = 40 KIAS

Figure 4-2.

4-4

CDP = 30 KIAS AT 15 FT (4.6m)

Rejected takeoff distance required (Sheet 1 of 3)


FAA APPROVED

REJECTED TAKE-OFF DISTANCE

REQUIRED

PART B

-40

Figure 4-2.

Rejectedtakeoff distance required (Sheet 2 of 3)

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS- 10.3

FAA APPROVED

REJECTED

TAKE-OFF DISTANCE

REQUIRED

PART C

GW 7.500 TO 11,900

VTOSS = 55 KIAS


0

OAT - °C

EXAMPLE:

OAT PRESSUREALTITUDE WIND

LB (3402 TO 5398 Kg)

50

0

=25°C -2000 FFT.(609. -30 KT!

FOLLOWARROWSFROM ( TO READ REJECTEDTAKE OFF DISTANCE REQUIRED = 580 FT. (177 m.) AT (@

FT. 200 600 1000 1400 1800 METERS 60 100 150 200 250 300 350 400 450 500 550 REJECTEDTAKE-OFF DISTANCE REQUIRED

Figure 4-2.

4-6


I

T N I R P N O R I O T / F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-10.3

FAA APPROVED

TAKE-OFF SPACE REQUIRED PART A

GW 7,500 TO 10.000 LB (3402 TO 4536 Kg)


VTOSS =40 KIAS

I

I

0

EXAMPLE:


=25°C =2000FT (609.6m) =20KTS

FOLLOW ARROWSFROM ( TO READ TAKE-OFFSPACE REQUIRED

= 535 FT. (163 m.) AT(

30 600 FT. 200 METERS 60 100 150 200 250 TAKE-OFF SPACEREQUIRED

Figure 4-3.

1000 300

Takeoff space required chart (Sheet 1 of 3)

4-7


BHT-412-FMS-10.3

TAKE-OFF SPACE REQUIRED PART B

GW 7.500 TO 10.800

VTOSS = 45 KIAS

LB (3402 TO 4899 Kg)

COP = 35 KIAS AT 20 FT (6.1m)

I I

Figure 4-3.

4-8


BHT-412-FMS-10.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TAKE-OFF SPACE REQUIRED PART C

Figure 4-3.



BHT-412-FMS-10.3

PT6T-3BE ENGINES TAKEOFF FLIGHT PATH 35 TO 500 FEET AGL PART A

2.5 MINUTE OEI POWER ENGINE RPM 97% GENERATOR 150 AMPS

Vtoss = 40 K IAS HEATER AND ANTI-ICE OFF INOPERATIVE ENGINE SECUFRED

uw

w La:

WI.-

-4

0

)

34TERS)

0)

MEAN HEIGHT GAINED IN 100 FEET (30.5 METERS) HORIZONTAL DISTANCE - FEET (METERS)

Figure 4-4.

4-10

Rev. 1

Takeoff flight path for PT6T-3BE engines (Sheet 1 of 6)


FAA APPROVED

PT6T-3BE ENGINES TAKEOFF FLIGHT PATH 500 TO 1000 FEET AGL PART A

I

Vtoss = 40 KIAS HEATER AND ANTI-ICE OFF INOPERATIVE ENGINE SECURED

30 MINUTE OEI POWER ENGINE RPM 97% GENERATOR 150 AMPS

4000

2000

MEAN HEIGHT

Figure 4-4.

GAINED IN 100 FEET(30.5

METERS)

HORIZONTAL

DISTANCE

- FEET (METERS)

I

Takeoff flight path for PT6T-3BE engines (Sheet 2 of 6)

Rev. 1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVE

BHT-412-FMS-10.3

PT6T-3BE ENGINES TAKEOFF FLIGHT PATH 35 TO 500 FEET AGL PART B

2.5 MINUTE OEI PC ENGINE RPM 97% GENERATOR 150 A

HEATERAND ANTI-ICE OFF

4000

-4200

I

-3800

-3600

3400

0

(0)

2

4

(1)

6

(2)

8

10

(3)

12

14

(4)

16

(5)

18

20

22

24

26

28

30

32

34

(6)

MEAN HEIGHT GAINED IN 100FEET (30.5 METERS)HORIZONTAL DISTANCE- FEET (METERS)

Figure 4-4. Takeoff flight path for PT6T-3BEengines (Sheet3 of 6) 4-12

Rev. 1


PT6T-3BE ENGINES TAKEOFF FLIGHT PATH 500 TO 1000 FEET AGL PART B


4000

HEATER AND ANTI-ICE OFF INOPERATIVE ENGINE SECURED

-1200

-1000

I

Figure 4-4. Takeoff flight path for PT6T-3BE engines (Sheet 4 of 6) Rev. 1

4-13

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 2.5 MINUTE OEI POWER ENGINE RPM 97% GENERATOR 150 AMPS 4000

Figure 4-4. Takeoff flight path for PT6T-3BE engines (Sheet 5 of 6)

4-14

Rev. 1


FAA APPROVED

I

PT6T-3BE ENGINES TAKEOFF FLIGHT PATH 500 TO 1000 FEET AGL PART C


Vtoss = 55 KIAS HEATER AND ANTI-ICE OFF INOPERATIVE ENGINE SECURED

LJOr

IL I-

4

I

MEAN HEIGHT

GAINED

IN 100 FEET (30.5 METERS) HORIZONTAL

DISTANCE - FEET (METERS)

I

Figure 4-4. Takeoff flight path for PT6T-3BE engines (Sheet 6 of 6) Rev. 1


FAA APPROVED

PT6T-3BG ENGINES TAKEOFF FLIGHT PATH 35 TO 500 FEET AGL PART A

ENGINERPM 97% GENERATOR150 AMPS

HEATER AND ANTI-ICE OFF INOPERATIVE ENGINE SECURED -1200

4000

OAT

-4600

MEAN HEIGHT GAINED IN 100FEET (30.5 METERS)HORIZONTAL DISTANCE - FEET (METERS)

Figure 4-4A.

4-16

Rev. 1

Takeoff flight path for PT6T-3BG engines (Sheet 1 of 3)


FAA APPROVED

PT6T-3BG ENGINES TAKEOFF FLIGHT PATH 35 TO 500 FEET AGL PART B 30 MINUTEOEI POWER Vtoss = 45 KIAS HEATER AND ANTI-ICE OFF ENGINERPM 97% INOPERATIVEENGINE SECURED GENERATOR150 AMPS

-1200

OAT

=

.20°C _

-

-30C

18

20

0

30

0

2

4

MEAN HEIGHT

Figure 4-4A.

6

6

10

12

14

16

22

GAINED IN 100 FEET (30.5 METERS) HORIZONTAL

24

26

28

30

32

34

DISTANCE - FEET (METERS)


Rev. 1

4-17


BHT-412-FMS-10.3

PT6T-3BG ENGINES TAKEOFF FLIGHT PATH 35 TO 500 FEET AGL PART C Vtoss = 55 KIAS 30 MINUTEOEI POWER HEATERAND ANTI-ICE OFF ENGINERPM 97% INOPERATIVE ENGINE SECURED GENERATOR150 AMPS

MEAN HEIGHT GAINED IN 100 FEET (30.5 METERS) HORIZONTAL DISTANCE- FEET (METERS)

Figure 4-4A.

4-18

Rev. 1



FAA APPROVED

SPACE REQUIRED

LANDING

-

50 FT. HEIGHT TO STOP

GW 7,500 TO 11,900 LB (3402 TO 5398 Kg)

DECISION POINT 100 FEET (30.5 METERS) @ 40 KIAS

LANDING

RATE OF DESCENT 500 FT/MIN.

I

40

.30

EXAMPLE:

.20

.10

I

10 ° OAT - C

0


-25°C =2000FT. (609.6m) =10 KTS

FOLLOWARROWSFROM TO READ LANDING SPACE REQUIRED = 420 FT (128M) AT

-

()

30

FT

200

400

600

800

METERS 100 150 200 250 LANDING SPACEREQUIRED

Figure 4-5.

Landing space required chart

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-10.4

SUPPLEMENT

GROUP"A" OPERATIONS 412-705-018

S/N 36087 AND SUB CAACERTIFIED 11 NOVEMBER 1996

This supplement shall be attached to Model 412EP Flight Manual when GROUP"A" OPERATIONS kit has been installed. Information contained herein supplements Information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

COPYRIGHT NOTICE COPYRIGHT 1996

11 NOVEMBER 1996

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-1 0.4

NOTICEPAGE

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 76101-0482


LOGOF REVISIONS

Original ...........

0 ..............

11 NOV 96

LOGOF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title .................................... NP ......................................

0 0

PAGE

A - B .................................. i- ii .................................. 1 - 34 ..................................

REVISION NO. 0 .0 0

NOTE


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O CAA APPROVED

BHT-412-FMS-CAA-10.4



0..............11

CIVIL AVIATION AUTHORITY SAFETY REGULATIONGROUP AVIATION HOUSE SOUTH AREA GATWICK AIRPORT GATWICK WEST SUSSEX RH6 OYR

NOV 96


GENERAL INFORMATION

INTRODUCTION

The information contained in this supplement is for group "A" operations. For limitations, normal procedures, emergency and malfunction procedures, and performance data not contained in this supplement, consult the appropriate sections of this flight manual. DEFINITIONS: GROUP "A"

TAKEOFF

Operation of the helicopter in such a manner that if

one engine fails at any time after the start of the

1. At or prior to CDP, return to and safely stop on the takeoff area; or 2. At or after CDP, climb out from point of failure and attain single engine forward flight.

GROUP"A" LANDING

Operation of the helicopter in such a manner that if one engine fails at any time after the start of a landing approach the helicopter can:

1. At or after LDP, continue the approach and safely land and stop on the clear heliport; or 2. At or prior to LDP, climb out from point of failure and attain single engine forward flight.

CRITICALDECISION POINT

The last point In the takeoff path at which a rejected takeoff can be assured, and the first point at which a completed takeoff can be assured.

LANDINGDECISION POINT

That point on the landing profile after which the helicopter is committed to landing

COMPLETED TAKEOFF DISTANCE REQUIRED

The horizontal distance from the start of the prescribed takeoff procedure to a point at least 50 and a feet above the takeoff surface where VTOSS positive rate of climb are attained following an engine failure occurring at or after CDP.

REJECTED TAKEOFF DISTANCE REQUIRED

The horizontal distance from the start of the prescribed takeoff procedure to the point where the helicopter is brought to a safe stop on the designated surface following an engine failure occurring at or prior to CDP.

TAKEOFF SAFETY PATH

The distance traveled from where the aircraft at or above 50 feet AGL to 1,000 feet reaches VTOSS AGL.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-10.4 TAKEOFF SAFETY

The airspeed that will assure the required climb performance with one engine inoperative

LANDING DISTANCE REQUIRED

The horizontal distance necessary to achieve a takeoff flight path at VTOSS and an altitude of 50 feet or higher, with one engine Inoperative at or prior to LDP; or the horizontal distance necessary to land the helicopter without further Incident, with one engine inoperative at or after LDP.

BALKED LANDING

The discontinuation of a landing approach and the initiation of a climbout. Group "A" balked landing capability following an engine failure is assured at or prior to LDP.

ABBREVIATIONS: AGL

Above Ground Level

CDP

Critical Decision Point

CDT

Critical Decision Time

GROC

Gross Rate of Climb

LDP

Landing Decision Point

VMI N IFR

Minimum Airspeed for

VTOSS (V 2 ) Vy

IFR

Takeoff Safety Speed Best Rate of Climb

Speed

WAT

Weight-Altitude-

Temperature

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-10.4

CAA APPROVED

1-5. CONFIGURATION

has not been conducted

Standard landing gear or high skid gear with or without emergency floats (floats stowed).

1-7. AIRSPEED

1-6. WEIGHT AND CENTER OF GRAVITY 1-6-C. TAKEOFFAND LANDING WEIGHTVS ALTITUDE

Refer to Gross Weight-Altitude-Ambient Air Temperature Limits Chart (figure 1-1). Charts designated Part A may be used for gross weights to 10,000 pounds (4636 kg). Part B charts may be used for gross weights to 10,800 pounds (4899 kg). Part C charts may be used for gross weights to 11,900 pounds (5398 kg).

Parts A, B, and C.

in areas between

1-7-A. CROSSWINDLIMITATIONS

The crosswind limit for takeoff and landing is 20 knots. Refer to the unfactored Chart in Component Headwind PERFORMANCEsubsection. Takeoff or landing downwind or with quartering tailwinds is prohibited.

1-8.

ALTITUDE

Maximum altitude for takeoff and landing is 4000 feet pressure altitude.

Interpolation of data between charts for different parts is not permitted. Testing

1


CAA APPROVED

PART A

W.A.T. LIMITS CHART

VTOSS IS 40 KIAS G.W. TO 10,000 LBS (4536 kg)

GROSSWEIGHT

PART

A

WEIGHT - ALTITUDE

- TEMPERATURE FOR TAKE-OFF

40 KIAS VTOSS G.W. TO 10,000 LB (436

AND I ANDING

kg)

412FS-CAA-10.4-1-1-1

Figure 1-1.

Gross weight-altitude-ambient air temperature landing (Sheet 1 of 3)

limits chart - takeoff and



PART B W.A.T. LIMITS CHART

VTOSS IS 45 KIAS G.W. TO 10.800 LBS (4899 kg)

6000

5000

3000

2000

1000

GROSS WEIGHT

PART

B

WEIGHT-

ALTITUDE - TEMPERATURE FOR TAKE-OFF

46 KIAS VTOSS G.W. TO 10.800

AND LANDING

LB (4899 kg)

412FS-CAA-10.4-1-1-2

Figure 1-1. Gross weight-altitude-amblent air temperature limits chart - takeoff and landing (Sheet 2 of 3) 3


CAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O PARTC

W.A.T.UMITS CHART

VTOSS IS 55 KIAS G.W. TO 11900 LBS (5398 kg)

412FS-CAA-10.4-1-1-3

Figure 1-1.

Gross weight-altitude-amblent

air temperature limits chart - takeoff and

landing (Sheet 3 of 3)


CAA APPROVED

2-7. BEFORE TAKEOFF

Perform Power Assurance Check, refer to basic manual.

2-8. TAKEOFF

2-8-A. STANDARDTYPE TAKEOFF

After attaining CDP, accelerate the helicopter to 65 KIAS and continue the climb.

Obtain CDP information - Refer to Limitations of this Supplement and figure 2-1. Collective ENG -


Flat pitch.

100%RPM (Nl).

Altimeter - Set, note indication collective fully down. Instruments -

with

Normal operating range.

SEAT BELT and NO SMOKEswitches - As desired. Area -

NOTE CDP height is determined by reference to the pilots barometric altimeter. Indicated altitude with collective full down on the takeoff surface is used as a ground level reference.

Clear.

Hover at approximately 4 feet (1.2 meters) skid height and note torque. Adjust ADI pitch bar to indicate level.

Initiate a takeoff from hover using a MAST TORQUE of °10% above that required to hover and 10 nose down attitude. NOTE Do not exceed MAST TORQUE, ITT, or GAS PROD RPM(N1) limits.

Maintain pitch attitude as the helicopter moves forward to achieve the correct CDP shown on the takeoff flight path profile diagram (figure 2-1).

2-10-A. STANDARDTYPE LANDING

NOTE type landing is A standard initiated from a LDP at 40 KIAS and an altitude of 100 feet above the runway, either in level flight or with a rate of decent of not more than 500 feet per minute (figure 22).

Flight controls level.

Adjust friction to desired

GOV switches -

AUTO.

Throttles -

Fully open.

ENG RPM -

100% RPM(N,,).

FORCETRIM switch - As desired STEP switch - As desired. Altimeter station.

-

Set to nearest reporting

SEAT BELT and NO SMOKE sign desired.

As 5

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TAKE-OFFFLIGHTPATHPROFILE MAXIMUMCONTINGENCY POWERRATINGAFTERPOWER FAILURE.THENREDUCETO MAXIMUM CONTINUOUSOEIAT 500 FT (152m) AGL.

HEIGHT

VTOSSMINIMUM

412FS-CAA-10.4-2-1

Figure 2-1.

Takeoff flight

path profile


CAA APPROVED

LANDING FLIGHT PATH PROFILE

FT.

DISTANCE LANDING

412FS-CAA-10.4-2-2

Figure 2-2.

Landing flight path profile

7


CAA APPROVED

Section3

3-1. INTRODUCTION

Tables 3-1 through 3-4 list panel wording, fault conditions, and corrective actions for emergencies and malfunctions that might occur during takeoff after CDP, during landing prior to LDP, and during landing after LDP.

whichever is higher. Climb out to 500 feet above the takeoff surface and accelerate to 65 KIAS. Reduce power to maximum continuous OEI limits. Refer to single engine enroute weight-altitude-temperature chart (figure 3-1). CAUTION

3-3. ENGINE

3-3-A. SINGLEENGINEFAILURE DURINGTAKEOFF PRIORTO CDP

An engine failure prior to reaching CDP will necessitate a landing back to the takeoff surface. If height permits, a positive deceleration to reduce forward airspeed is required. As the helicopter descends, it should be leveled and the collective should be used as required to cushion the landing. Some forward ground speed is normally required at touchdown. Maintain control of the helicopter.

Collective - Adjust to maintain ROTOR RPM and OEI power limits. Flair to reduce ground speed.


attitude

before


Complete shutdown of affected engine. DURINGTAKEOFF AFTER CDP

In the event of an engine failure following CDP, airspeed should be Increased to the takeoff safety speed (TOSS)or maintained,

DURING COLD WEATHER OPERATIONS, CAREFULLY MONITOR TORQUE OF THE NORMAL ENGINE WHEN ONE ENGINEFAILS OR IS SHUT DOWN IN FLIGHT. NOTE

During takeoff, after CDP, it is permissible to droop ROTOR RPM to 91% during the transition from twin engine to single engine flight an engine failure. following ROTOR RPM should be regained to normal operating range at or before attaining appropriate best rate of climb speed.


Collective - Adjust to maintain ROTOR RPM and OEI power limits. Airspeed

-

If below

VTOSS, smoothly

increase to VTOSS and initiate a climb. Throttle (affected engine) - Close.

Complete shutdown of affected engine. ENG (unaffected engine) RPM (Nll). AFTER CDP -

Set to 100%

ENROUTECLIMB


CAA APPROVED

Refer to figure 3-2 Enroute net rate of climb chart.

* DURINGLANDING PRIORTO LDP


Collective

-

Adjust to maintain ROTOR

RPM and OEI power limits. Flair to reduce speed.

Execute the same procedures as for single engine failure on takeoff after CDP or proceed to LDP and use the procedure below.


* DURINGLANDING AFTER THE LDP


attitude

before


The helicopter, with an emergency, is committed to land after LDP. The landing is accomplished using up to the maximum power of the remaining engine while maintaining rotor speed within limits.

9


CAA APPROVED

6000

4000

3000

2000

i

-

1000

78

80

36

82

37

84

86

88

38

39

40

90

41

92

94

96

98

42 43 44 GROSS WEIGHT

100

45

102

46

104

47

106

108

110

112

48

49

5'0

51

114

116

52

53

119 54

ENROUTEWEIGHT- ALTITUDE- TEMPERATURE- MAXIMUMCONTINUOUSOEI POWER - VY = 70 KIAS 7,500TO 11,900 LB (3402 TO 5398 Kg)

412FS-CAA-10.4-3-1

Figure 3-1. Single engine enroute weight-altitude-temperature chart



10,000

12,000 (5443) 11,500\

(5216)

11,000 (4990) 10,500 (4763) 10,000 (4536)

600

2000

NET R/C - FT.MIN

ENROUTENET RATEOF CLIMB - ONE ENGINEINOPERATIVE MAXIMUMCONTINUOUSOEI - 70 KIAS - ENGINE RPM 97% GENERATOR150AMPS - HEATEROFF - INOP ENGINESECURED

Figure 3-2.

412FS-CAA-10.4-3-2

Enroute net rate of climb chart

11


CAA APPROVED

Table 3-1. Warning lights-Takeoff

PANEL WORDING

prior to CDP

FAULT CONDITION

CORRECTIVE ACTION

MASTER CAUTION


Land Immediately

FIRE PULL (1 OR 2)

Fire indication in No. 1 or Land immediately. No. 2 engine compartment Pull affected FIRE PULL handle. Select MAIN fire extinguisher; it necessary, select RESERVE fire extinguisher.

BAGGAGE FIRE


ENG OUT (1 or 2)

GAS PRODabnormally low, Land immediately. below 53 + 2% RPM (N,), on Refer to ENGINEOUT procedures. No. 1 or No. 2 engine.

X MSN PRESSURE


XMSN OIL TEMP C BOX PRESSURE

baggage





C BOX TEMP

Combining gearbox temperature above limit.

Land immediately.

BATTERY TEMP


Land immediately. BATTERY BUS 1 and BUS 2

switches -

OFF.

WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER

ILLUMINATION OF BATTERY TEMP

LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS PRIORTO RETURNTO SERVICE.

ROTOR BRAKE


not

Land immediately.


CAA APPROVED


PANEL WORDING

after CDP

FAULT CONDITION

CORRECTIVE ACTION

MASTERCAUTION

Warning or Caution light(s) illuminated

Accelerate

FIRE PULL

Fire indication in No.1 or Accelerate to VToss. Pull affected No. 2 engine compartment. FIRE PULL handle. Select MAIN fire extinguisher. Close throttle of affected engine. Select RESERVE fire extinguisher if necessary. Land as soon as possible.

(1 or 2)

to VTOSS.Reset MASTER

CAUTION light; take appropriate corrective action as required by illuminated segment(s).

BAGGAGEFIRE


ENG OUT (1 or 2)

GAS PROD abnormally low, Accelerate to VTOSS. Secure below 53 ± 2% RPM (N,), on appropriate engine. Refer to ENGINE No.1 or No.2 engine OUT procedure. Land as soon as possible.

X MSN PRESSURE

C BOX TEMP


O I L Transmission oil pressure Accelerate to VTOSS.Reduce power; below limit. verify fault on XMSN OIL pressure gage. Land as soon as possible.

XMSNOIL TEMIP

C BOX PRESSURE

baggage


Accelerate

to VTOSS.Reduce power;

verify fault on XMSN OIL temperature gage. Land as soon as possible.


Accelerate to VTOSS. Reduce power; verify fault on GEAR BOX pressure gage. Land Immediately


Accelerate to VTOSS. Reduce power; verify fault on GEAR BOX temperature gage. Land as soon as possible.

13


CAA APPROVED


PANEL WORDING

BATTERYTEMP

after CDP (Cont)

FAULT CONDITION

CORRECTIVE ACTION

Battery case temperature above limit

Accelerate to VTOSS.BATTERY BUS 1 and BUS 2 switches - OFF. Land as soon as practical. WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATION OF BATTERY

TEMP

LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURNTO SERVICE.

ROTORBRAKE

Rotor brake linings retracted

not

Accelerate

to VTOSS. Check

rotor

brake handle fully up in detent. If light remains on, land as soon as possible.


CAA APPROVED

Table 3-3. Warning lights-Landing

PANEL WORDING

prior to LDP

FAULT CONDITION

CORRECTIVE ACTION

MASTER CAUTION


Reset MASTER CAUTION light; take appropriate corrective action as required by illuminated segment.

FIRE PULL (1 or 2)

Fire indication in No.1 or Pull affected FIRE PULL handle. No.2 engine compartment. Select MAIN fire extinguisher. Close throttle of affected engine. Select RESERVE fire extinguisher, if necessary. Land as soon as possible

BAGGAGE FIRE

Smoke in baggage compartment.

ENG OUT (1 or 2)

GAS PROD abnormally

XMSN PRESSURE

low, Maintain VTOSS.Secure appropriate below 53 2% RPM (N1), on engine. Land as soon as possible. No.1 or No.2 engine. Refer to ENGINE OUT procedure.


XMSN OIL TEMP

C

BOX

PRESSURE



Reduce power. Verify fault on XMSN OIL pressure gage. Land immediately. Reduce power. Verify fault on XMSN OIL temperature gage. Land as soon as possible.


Reduce power. Verify fault on GEAR BOX pressure gage. Land as soon as possible


Reduce power. Verify fault on GEAR BOX temperature gage. Land as soon as possible.

C BOX OIL TEMP

15


CAA APPROVED


PANEL WORDING

BATTERYTEMP

prior to LDP (Cont)

FAULT CONDITION

CORRECTIVE ACTION


BATTERY switches practical

BUS 1 and BUS 2 OFF. Land as soon as

WARNING

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATIONOF BATTERY TEMP LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS PRIORTO RETURNTO SERVICE.

ROTOR BRAKE


not

Check rotor brake handle fully up In detent. If light remains on, land as soon as possible.


CAA APPROVED


PANEL WORDING

after LDP

FAULT CONDITION

CORRECTIVE ACTION

MASTER CAUTION


Land immediately

FIRE PULL (1 or 2)

Fire indication in No.1 or Land immediately. No.2 engine compartment. Pull affected FIRE PULL handle. Close throttle of affected engine. Select MAIN fire extinguisher; if necessary, select RESERVE fire extinguisher.

BAGGAGEFIRE

Smoke in compartment.

ENG OUT (1 or 2)

GAS PROD abnormally low, Land immediately. below 53 ± 2% rpm (N), on Refer to ENGINEOUT procedures. No.1 or No.2 engine.

X MSN

PRESSURE


XMSN OIL TEM P C BOX PRESSURE

C BOX TEMP

BATTERY TEMI

baggage


O I L Combining gearbox oil pressure below normal

Land immediately. Inspect tailboom area for damage


Combining gearbox temperature above limit.

Land immediately.

Battery case temperature

Land immediately. BATTERY BUS 1 and

above limits.

switches -

BUS 2

OFF.

WARNING

BATTERY

SHALL

FOR ENGINE

ILLUMINATION

NOT BE USED

START

AFTER

OF BATTERY

TEMP

LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURNTO SERVICE.

ROTORBRAKE

Rotor

brake

retracted.

linings

not

Land immediately.

17

T N I Section4 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-10.4

4-1. INTRODUCTION

The power performance data presented in this section is based on engine manufacturers minimum specification power for the PT6T-3D engine with installation losses.

The takeoff and landing data presented in this section is based on test performed on a level asphalt runway 75 feet wide. The minimum runway length for standard takeoff and landing procedures varies with wind, gross weight, pressure altitude, and temperature.


Refer to Section 4 of applicable Basic Flight Manual for power assurance charts to determine if the engine (power sections) can produce installed specification power.

The hover check is preformed prior to takeoff. The in-flight check is provided for in-flight monitoring of engine performance. If either engine (power section) does not meet the requirements of the hover or inflight power assurance check, group "A" performance will not be achievable. The cause of engine power loss, or excessive Interturbine temperature (ITT) or GAS PROD RPM (N,) shall be determined as soon as practical. Refer to appropriate engine maintenance manual.

4-5. HOVER CEILING 4-5-A. HEADWINDCOMPONENT

CAA APPROVED

The Unfactored Headwind Component chart (figure 4-1) is provided with an example to determine critical crosswind and corrected headwind for group "A" takeoff and landings. The headwind component, as calculated from the headwind component chart, is applied to parts A, B, and C of the Takeoff Space Required charts.

4-6. TAKEOFF DISTANCE 4-6-A. REJECTEDTAKEOFF DISTANCEREQUIRED

The rejected takeoff distance required Is the space necessary to takeoff, climb to CDP, encounter an engine failure at CDP, return to takeoff surface, and stop safely. The rejected takeoff distance required is obtained from either part A, B, or C of the Rejected Takeoff Distance Required charts (figure 4-2).

4-6-B. TAKEOFFSPACE REQUIRED

The takeoff space required is the horizontal distance required to takeoff, climb to CDP, encounter an engine failure, accelerate to VTOSS,and climb to 50 feet AGL above the takeoff space. Takeoff space required is obtained from either part A, B, or C of the Takeoff Space Required charts (figure 4-3) using the headwind component from the Unfactored Headwind Component chart (figure 4-1).


4-6-C. TAKEOFF FLIGHT PATH

The takeoff flight path begins at the end of Takeoff Space Required, at 50 feet AGL or higher, above the takeoff space and at VTOSS,Parts A, B, or C of the Takeoff Flight

Path charts (figure 4-4) provide data for 50 to 500 feet and 500 to 1,000 feet AGL. These charts provide altitude gain for each 100 feet (30.5 meters) horizontal distance traveled.

4-7. CLIMB AND DESCENT 4-7-A. ENROUTECLIMB

Refer to figure 4-5 OEI Enroute net climb gradient chart.


Landing space required is the distance necessary to come to a stop from LDP with one engine inoperative.

Landing space required is obtained from the Landing Space Required chart (figure 4-6).

4-9-B. CORRECTED LANDING DISTANCE

Corrected landing distance from LDP is landing distance corrected for wind factor.

The headwind component is obtained from calculation of the Unfactored Headwind Component chart (figure 4-1), and applied to the Landing Space Required chart (figure 4-6) to obtain corrected landing distance.

4-9. LANDING DISTANCE 4-9-A. LANDINGSPACE REQUIRED

19


CAA APPROVED

40

x

35

35 30 25

CROSSWIND COMPONENT-

KT.

EXAMPLE

4.

REPORTED

6.

FOLLOW

WIND

SPEED

THE SHAPE

COMPONENT

SCALE

...............................

OF THE

AND

CURVED

READ

UNES.

HEADWIND

20 KNOTS

TO WIND

COMPONENT

.......

17

KNOTS

412FS-CAA-10.4-4-1

Figure 4-1.

Unfactored headwind component chart


CAA APPROVED

REJECTED TAKE-OFF AREA REQUIRED PART A

GW 7.500 TO 10.000 LB (3402 TO 4536 Kg)

VTOSS = 40 KIAS


10

OAT - C

Figure 4-2.


21



REJECTED TAKE-OFFAREA REQUIRED PART B

GW 7,500 TO 10.800 LB (3402 TO 4899 Kg)

Figure 4-2.

22


CAA APPROVED


REJECTED TAKE-OFF AREA REQUIRED PART C

I

0

EXAMPLE:

10

20

OAT PRESSURE ALTITUDE WIND

REJECTED TAKE-OFF AREA REQUIRED

412FS-CAA-10.4-4-2-3

Figure 4-2.


23


CAA APPROVED

TAKE-OFF SPACE REQUIRED PART A

TAKE-OFFSPACEREQUIRED

412FS-CAA-10.4-4-3-1

Figure 4-3.




TAKE-OFF SPACE REQUIRED PART B

GW 7.500 TO 10.800 LB (3402 TO 4899 Kg)

VTOSS = 4

EXAMPLE:

KIAS


OA PR WI FO TO

RE

Figure 4-3.


25


CAA APPROVED

TAKE-OFF SPACE REQUIRED PART C

GW 7,500

TO 11,900

VTOSS = 56 KIAS

LB (3402

TO 5398

Kg)


TAKE-OFF SPACE REQUIRED

412FS-CAA-I0.4-4-3-3 412FS-CAA-10.4-4-3-3

Figure 4-3. Takeoff space required chart (Sheet 3 of 3)

0



412FS-CAA-10.4-4-4-1

Figure 4-4.

Takeoff flight path (Sheet

1 of 6)

27


CAA APPROVED

VTOSS = 40 KIAS

PART A

Figure 4-4.

Takeoff flight

path (Sheet 2 of 6)


CAA APPROVED

PART B

2000 (609.6)

10.800

10,000 (4536)

MEAN HEIGHT GAINED IN 100FEETHORIZONTAL DISTANCE- FEET (METERS) 45 KNOTS -

MAXIMUM

CONTINGENCY

POWER

412FS-CAA-10.4-4-4-3

Figure 4-4.

Takeoff flight path (Sheet 3 of 6)


CAA APPROVED

MEAN HEIGHT GAINED IN 100 FEET HORIZONTAL DISTANCE - FEET (METERS) 45 KNOTS- MAXIMUM CONTINUOUSOEI 500 TO 1000 FEET (152 TO 305 METERS) ABOVE SURFACE

412FS-CAA-10.4-4-4-4

Figure 4-4. Takeoff flight path (Sheet 4 of 6)



PART C

I

c

10

30

41 2FS-AA-10.4-4-4-5

Figure 4-4.

Takeoff flight path (Sheet 5 of 6)

31


CAA APPROVED

PART C

OAT-'C

MAX. GROSS

(3402)

Figure 4-4. Takeoff flight path (Sheet 6 of 6)


2000

ENROUTENET GRADIENTOF CLIMB - ONE ENGINE INOPERATIVE MAXIMUMCONTINUOUSOEI - 70 KIAS - ENGINE RPM97% GENERATOR150 AMPS - HEATEROFF - INOP ENGINESECURED

412FS-CAA-10.4-4-5

Figure 4-5.

OEI Enroute net climb gradient chart

33


CAA APPROVED

LANDING SPACE REQUIRED -

100 FT. HEIGHT TO STOP

GW 7,500 TO 11,900 LB (3402 TO 5398 kg)

LANDING

DECISION POINT 100 FEET (30.5 METERS) @ 40 KIAS RATE OF DESCENT 500 FT/MIN.Figure

40

-30

-20

-10

0

10 OAT - C

Figure 4-6.

20

I ® 30

40

50

Landing space required chart

4-6.Landing

space

required

T N I R 3ell Model 412 P N O R I O T ROTORCRAFT F C U T FLIGHT O OD MANUAL N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-11.1

33001 THROUGH33107

SUPPLEMENTFOR 20/20 AUXILIARYFUELOPERATIONS (412-899-139)

CERTIFIED JANUARY 29, 1982

This supplement shall be attached to the Model 412 Flight Manual (S/N 33001 - 33107) when the 412-899-139 20/20 Auxiliary Fuel Kit has beeninstalled. The information contained hereinsupplements the information of the basic Flight Manual. For Limitations, Procedures, and PerformanceData not contained in this supplement consult the

A Subsidiary of TextronInc

POST OFFICE BOX 482 * FORT WORTH. TEXAS 76101

COPYRIGHT1989 BELLxHELICOPTER TEXTRON INC. ALL RIGHTS RESERVED

29

JANUARY 1982 REISSUED 8 MAY 1989 REISSUED 8 MAY 1989

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Log of Pages


BHT-412-FMS-11.1


LOG OF REVISIONS

Original ..... Reissued ....

0 ........ 0 ...........

.October 2, 1981 May 8, 1989

LOGOF PAGES

Page

Tide ....................... A......................... i/ii......................... 1-6...................... 7/8........................

Revision No. 0 0 0 0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONDIRECTORATE AIRCRAFTCERTIFICATION SERVICE DEPARTMENTOF TRANSPORTATION SOUTHWESTREGION,FORTWORTH, TEXAS


A

Useor disclosure of data containedonthis pageis subjectto the restrictionon the title pageof this document.



INTRODUCTION

BHT-412-FMS-11.1

INTRODUCTION

The 20/20 Auxiliary Fuel Kit provides additional fuel capacity to extend the range of the helicopter. The kit includes two twenty gallon fuel tanks, a fuel quantity gage, and the plumbing necessaryto connect them to the basic helicopter fuel system. The tanks are mounted in the aft cabin, one on each side of the pylon structure. Cushionsare installed on the tanks to allow passengerseating on the tanks.

Useor disclosure of datacontainedonthis pageissubjectto the restrictionon the tidepageof this document.

i/ii

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHTMANUAL


Section 1

BHT-412-FMS-11.1

NOTE

The contents of this supplement shall be used in conjunction with the basic Flight Manual for helicopters equipped with the 20/20 Auxiliary Fuel Tank Kit installed.

WEIGHTAND BALANCE

Actual weight change shall be determinedafter kit is Installed and ballast readjusted. if necessary,to retain empty weight CG within allowable limits.

PLACARDSAND DECALS

FUEL FUELSYSTEMUSABLECAPACITY BASIC AIRCRAFT211 U.S. GAL. = 798.6 LITERS WITH 412-706-007 AUX. KIT 374 U.S. GAL. = 1415.5 LITERS WITH 412-899-139 AUX. KIT 251 U.S. GAL. = 950.0 LITERS FUELTYPEPERPWASPEC522 SEEFLIGHTMANUAL FORAPPROVEDFUEL

Useor disclosureof datacontainedon this pageissubjectto the restriction n the title pageof this document.

1

Sections2. 3. 4 & 6



T N Section 2 I R P N O R 0 I O T F C Section 3 U T O OD N R E P E E R R C S N G O N I R W 0 O E FSection 5VI Y L 0 N O BHT-412-FMS-11.1

No Change

No Change

No Change

2

Useor disclosure of datacontainedon this pageis subjectto the restrictiononthe title pageof this document.

T N * Section 6 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

412 ROTORCRAFT FLIGH1MANUAL

Section 6

BHT-412-FMS-11.1

FUELLOADING

Total usable fuel capacity with 412-899-139 Auxiliary Fuel Kit installed is 251 U.S. gallons. Table6-1 lists fuel quantities, weights, centers of gravity, and moments for fuel quantities from 10 to 260 U.S. gallons. Moments are given In Inch-pounds.

Table 6-2 provides liters, weights in kilograms, centers of gravity in millimeters, and moments in kilogram millimeters divided by 100 for fuel quantities from 38 to 984 liters.

Useor disclosureof datacontainedon this pageis subjectto the restrictiononthe tite pageof this document.

3

T N I 0 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O Section 6



BHT-412-FMS- 1.1

Table 6-1.

Usable Fuel LoadingTable (English) (Sheet 1 of 2)

USABLE FUELLOADING TABLE

JET A. JET A-1. JP-5 - 6.8 LB/GAL

GAL.

WEIGHT LB.

C.G. STA.

MOMENT IN. LB.

10 20 30 40 50 60 70 '72.6 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260

68 136 204 272 340 408 476 494 544 612 680 748 816 884 952 1020 1088 1156 1224 1292 1360 1428 1496 1564 1632 1700 1768

143.6 143.6 140.2 134.8 131.6 129.4 127.9 127.6 128.3 130.9 134.0 136.3 138.3 140.0 141.3 142.6 143.7 144.7 145.7 146.7 147.7 148.6 149.5 150.4 151.2 151.9 152.6

9765 19530 28601 36666 44744 52795 60880 62994 69795 80111 91120 101952 112853 123760 134518 145452 156346 167273 178337 189536 200872 212201 223652 235226 246758 258230 269797

U.S.

* Most Forward Fuel C.G.

0

4




T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section 6

Table6-1. UsableFuelLoding Table(English)(Sheet2 of 21

USABLEFUELLOADINGTABLE JET B, JP-4 - 6.5 LB/GAL

U.S. GAL.

WEIGHT LB.

C.G. STA.

MOMENT IN. LB.

10 20 30 40 50 60 70 *72.6 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260

65 130 195 260 325 390 455 472 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1365 1430 1495 1560 1625 1690

143.6 143.6 140.2 134.8 131.6 129.4 127.9 127.6 128.3 130.9 134.0 136.3 138.3 140.0 141.3 142.6 143.7 144.7 145.7 146.7 147.7 148.6 149.5 150.4 151.2 151.9 152.6

9334 18668 27348 35055 42771 50480 58186 60176 66706 76577 87100 97455 107874 118300 128583 139035 149448 159894 170469 181175 192010 202839 213785 224848 235872 246838 257894

*Most Forward Fuel C.G.

Useor disclosureof datacontainedon this pageissubjectto the restrictiononthe title pageof this document. l

BHT-412-FMS-11.1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O 412 ROTORCRAFT FLIGHTMANUAL

Section 6


BHT-412-FMS-11.1

Table6-2. UsableFuelLoadingTable(Metric)(Sheet1 of 2) USABLEFUELLOADINGTABLE

JET A, JET A-1, JP-5 - 0.815 kg/1

LITERS 38 76 114 151 189 227 265 *275 303 341 379 416 454 492 530 568 606 643 681 719 757 794 833 871 908 946 984

WEIGHT C.G. MOMENT KILOGRAMS MILLIMETERS kg. mm. 31.0 61.9 92.9 123.1 154.0 185.0 216.0 224.1 246.9 277.9 308.9 339.0 370.0 401.0 431.9 462.9 493.9 524.0 555.0 586.0 616.9 647.1 678.9 709.9 740.0 771.0 801.9

3647 3647 3561 3424 3343 3287 3249 3241 3259 3325 3404 3462 3513 3556 3589 3622 3650 3675 3701 3726 3752 3774 3797 3820 3840 3858 3876

113057 225749 330817 421494 514822 608095 701784 726308 804647 924018 1051496 1173618 1299810 1425956 1550089 1676624 1802735 1925700 2054055 2183436 2314609 2442155 2577783 2711818 2841600 2974518 3108164

* Most Forward Fuel C.G.

6

Useor disclosureof datacontainedon this pageissubjectto the restrictiononthe title pageof this document.



Section 6

BHT-412-FMS-11.1

Table6-2. UsableFuelLoadingTable(Metric)(Sheet2 of 21

USABLEFUELLOADINGTABLE JET B. JP-4 - 0.779 kg/1

LITERS 38 76 114 151 189 227 265 *275 303 341 379 416 454 492 530 568 606 643 681 719 757 794 833 871 908 946 984

WEIGHT C.G. MOMENT KILOGRAMS MILLIMETERS kg. mm. 29.5 59.0 88.5 117.9 147.4 176.9 206.4 214.1 235.9 265.4 294.8 324.3 353.8 383.3 412.8 442.3 471.7 501.2 530.7 560.2 589.7 619.2 648.6 678.1 707.6 737.1 766.6

3647 3647 3561 3424 3343 3287 3249 3241 3259 3325 3404 3462 3513 3556 3589 3622 3650 3675 3701 3726 3752 3774 3797 3820 3840 3858 3876

107598 215198 315156 403680 492705 581428 670523 693907 768756 882418 1003381 1122733 1242836 1363015 1481548 1602028 1721696 1842101 1964004 2087406 2212307 2337133 2462929 2590451 2717524 2843923 2971372

*Most Forward Fuel C.G.

Useor disclosureof datacontainedonthis pageis subjectto the restrictiononthe title pageof this document.

7/8

T N I R P N i3el Madel 412 IO R O T F C ROTORCRAFT U T D MANUAL N O OFLIGHT R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-12

·

SUPPLEMENTFOR NIGHTSUN SEARCHLIGHT (212-899-333)

CERTIFIED DECEMBER4, 1981

This supplement shall be attached to the Model 412 Flight Manual when the 212-899-333 Nightsun Searchlight has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopters

A Subsidialy of TextronInc POST OFFICE BOX 482 * FORT WORTH, TEXAS 76101

COPYRIGHT"1989

BELL® HELICOPTER TEXTRON INC. ALL RIGHTS RESERVED

4 DECEMBER 1981 REISSUED8 MAY 1989


Log of Pages


BHT-412-FMS-12

LOG OF REVISIONS

Original ..... Reissued ....

0

........ ...........

December4, 1981 May 8, 1989

LOG OF PAGES

Revision No.

Page

Title - A ................... i/ii......................... 1-2...................... 3/4........................

Page

Revision No.

0 0 0 0

APPROVED:

MANAGER

DIRECTORATE ROTORCRAFTCERTIFICATION AIRCRAFTCERTIFICATIONSERVICE DEPARTMENTOF TRANSPORTATION SOUTHWESTREGION,FORTWORTH, TEXAS

NOTE: Revised text is Indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.




Introduction

BHT-412-FMS-12

INTRODUCTION

The Nightsun Searchlight is a high intensity light which mounts on the lower nose of the helicopter. The xenon arc light may be started, aimed, and focused from the operator's panel inside the helicopter. The kit consists of the Nightsun Searchlight, mounts, hardware, cable, and operator's panel.

Useor disclosureof datacontainedonthis pageis subjectto the restrictionon thetitle pageof this document.

T N I Section R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL


Sections 1

BHT-412-FMS-12

necessary, to return empty weight CG within allowable limits.

OPERATING LIMITATIONS

IFR operation is prohibited with Searchlight installed.

Nightsun

PLACARDS AND DECALS

FLIGHT CREW LIMITATIONS

Operation of the Nightsun Searchlight is restricted to the copilot or operator position.


Actual weight changes shall be determined after searchlight is installed and ballast readjusted, if

Useor disclosureof datacontainedonthis pageis subjectto the restrictiononthe title pageof this document.

CAUTION

DO NOT USE NIGHTSUN SEARCHLIGHT BELOW 50 FT AGL OR IN FOG CONDITIONS. MONITOR LOADMETER WHEN USING NIGHTSUN SEARCHLIGHT. 412099-5

T N I R Section 2 P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section 2

412 ROTORCRAFT FUGHT MANUAL

FAAAPPROVED SUPPLEMENT

BHT-412-FMS-12

EXTERIOR CHECK

Aim and focus - As desired.

Nightsun Searchlight - Security and wiring. Lens for cleanliness.

HOLDING SWITCH IN START POSITION AFTER IGNITION MAY DAMAGE EQUIPMENT.

PRESTART CHECK

SCHLT PWR and SCHLT CONT circuit breakers - IN.

INFLIGHT OPERATION

NIGHTSUN SEARCHLIGHT MASTER switch - ON.

NIGHTSUN SEARCHLIGHTSTART switch START, hold in start position approximately 5 seconds, or until ignition hasoccurred.

2

CAUTION

DO NOT AIM THE BEAM TOWARD OTHER AIRCRAFT OR VEHICLES BECAUSE OF TEMPORARY BLINDING EFFECT.

BEFORE LANDING

NIGHTSUN SEARCHLIGHT switch - OFF.

MASTER

Useor disclosureof datacontainedonthis pageissubjectto the restrictiononthe title pageof this document.

T N I R P N O R I O T F C U T D O Section O N 4 R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT


Sections3 & 4

BHT-412-FMS-12


No change from the basic Flight Manual.

Useor disclosureof data containedonthis pageis subjectto the restrictionon the title pageof this document.


SUPPLEMENTFOR FIXEDSTEP (212-706-057)

CERTIFIED 6 FEBRUARY1982

This supplement shall be attached to the Models 412 and 412 EP Flight Manual when the 412-706-057 fixed step has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter


NOTICE COPYRIGHT 1994 COPYRIGHT BELLHELICOPTER TEXTRON INC TEXTRON, HELICOPTER ANDBELL A DIVISION OFTEXTRON CANADA LTD ALLRIGHTS RESERVED

POST OFFICE BOX 482 * FORT WORTH, TEXAS76101

REISSUE -

23 JUNE 1994




BHT-412-FMS-15

LOG OF REVISIONS

Original ..... Reissue..... Reissue....

0 0 0

............ ............. .............

06 Feb 82 22 Feb 89 23Jun94

LOG OF PAGES

Page

Revision No.

Title....................... A......................... i/ii......................... 1 -2......................

Page

0 0 0 0

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ROTORCRAFTCERTIFICATIONOFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

NOTE: Revisedtext is indicated by a black vertical line. Insert latest revision pages; disposeof supersededpages.

Revision No.



INTRODUCTION

BHT-412-FMS-15

INTRODUCTION

The fixed steps mount to the sides of the fuselageto facilitate passengerentry and exit.

i/ii


SECTION1


BHT-412-FMS-15

OPERATINGLIMITATIONS

The contents of this supplement shall be used in conjunction with the basic Flight Manual for helicopters equippedwith the fixed step. The 412-706-004 Emergency Float Kit shall not be installedin conjunction with the fixed step.

The 212-706-105 Passenger Step shall not be installed in conjunction with the fixed step.

The 212-706-057 Fixed Step shall be removed when the 214-706-003 Internal Hoist is installed.

WEIGHT - CG LIMITATIONS

Actual weight change shall be determinedafter kit is installed and ballast readjusted, if necessary,to returnempty weight CGwithin allowable limits.

T N I Section 2 R P N O R I O T F C U T OSectionO3 D N R E P E E R R C S G Section 4ON N I R W O E I F V NLY O SECTIONS2. 3 AND 4



BHT-412-FMS-15

~


I

No changefrom basic Flight Manual.


T N I R P N ROTORCRAFT O R I FLIGHT MANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-17.2, 17.3, AND 17.4

Bell MODELS

SUPPLEMENT

AUXILIARY FUEL SUPPLEMENT 412-706-007

33108 - 33213 36001 - 36019 AND 36020 - 36086 AND 36087 AND SUB

CERTIFIED 5 JANUARY 1984

This supplement shall be attached to Bell Helicopter Model 412 and 412EP Flight Manuals when 412-706-007 Auxiliary Fuel Kit has been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter COPYRIGHT NOTICE 1994 COPYRIGHT COPYRIGHT 1994 BELL® HELICOPTER INC. ANDBELLHELICOPTER TEXTRON. A DIVISIONOF TEXTRONCANADALTD ALLRIGHTS RESERVED

A Subsidiary of Textron Inc POST OFFICE BOX 482 . FORT WORTH TEXAS 76101

REISSUE -

23 JUNE 1994

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-17.2, 17.3, AND 17.4

FAA APPROVED

NOTICEPAGE


Manufacturer's Data portion of this supplement is proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron, Inc.



BHT-412-FMS-17.2, 17.3, AND 17.4

LOG OF REVISIONS


Revision ..........

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May 22, 1987

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May 9, 1991

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MANUFACTURER'S DATA

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NOTE


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-17.2, 17.3, AND 17.4

FAA APPROVED

LOG OF APPROVED

Original ........... Revision ..........

0 ........ January 5, 1984 1 ........... May 22, 1987

APPROVED:

MANAGER


B

REVISIONS

Reissue ........... Reissue...........0

............ May 9,1991 .......... June 23, 1994


INTRODUCTION


BHT-412-FMS-17.2, 17.3 AND 17.4

INTRODUCTION

The Auxiliary Fuel Kit provides additional fuel capacity to extend the range of the helicopter. The kit consists of a left and right auxiliary fuel tank and the hardware and wiring necessary to complete the installation. The left or right auxiliary fuel tank may be removed as operational requirements dictate. One fuel tank provides an additional 81.7 U.S. gallons (309.2 liters) of fuel. Both fuel tanks combined. provide an additional 163.4 U.S. gallons (618.5 liters) of fuel.

i/ii

T N I * Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL


Section 1

BHT-412-FMS-17.2,17.3 AND 17.4

WEIGHT/CG

LIMITATIONS

Actual weight changes shall be determined after installation of auxiliary fuel tanks(s), and ballast shall be readjusted, if necessary, to return empty weight CG within allowable limits.

The gross weight center of gravity limits as presented in the basic Flight Manual do not change when either or both auxiliary fuel tanks are installed. Refer to Manufacturer's Data, Section 1, for weight and balance data and loading example.

WARNING

INDISCRIMINATE LOADING OF THE IN MAY RESULT HELICOPTER VIOLATION OF THE PERMISSIBLE CENTER OF GRAVITY LIMITATIONS WHEN THE HELICOPTER IS EQUIPPED AUXILIARY WITH THE 412-706-007 FUEL KIT.

1-1/1-2



Section 2

BHT-412-FMS-17.2.17.3 AND 17.4

IN-FLIGHT OPERATION CAUTION

WHEN ONLY ONE CABIN MOUNTED AUXILIARY FUEL TANK IS USED, THE FUEL INTCON SWITCH MUST BE REPLACED TO OPEN WHEN A FUEL QUANTITY INDICATION OF 500 LBS. IS ON EITHER SIDE. FAILURE TO MANUALLY OPEN THE FUEL INTCON SWITCH WILL RESULT IN FUEL EXHAUSTION TO THE ENGINE OPPOSITE THE SIDE WHICH HAS THE AUXILIARY TANK. THE FEATURE WHICH AUTOMATICALLY OPENS THE INTERCONNECTVALVE WILL NOT HAVE A CHANCE TO FUNCTION WITH ONLY ONE AUXILIARY TANK INSTALLED. WITH TWO EQUALLY LOADED AUXILIARY TANKS INSTALLED, THE AUTOMATIC FEATURE WILL FUNCTION.

2-1/2-2


412 ROTORCRAFT FLIGHT MANUAL 33108 CONFIGURATION

Section 3

BHT-412-FMS-17.2, 17.3AND 17.4

No change from

basic Flight Manual.

3-1/3-2


412 ROTORCRAFT FLIGHT MANUAL 33108 CONFIGURATION

Section 4

BHT-412-FMS-17.2, 17.3 AND 17.4


4-1/4-2

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT MANUFACTURER'S DATA

Section 1

BHT-412-FMS-17.2. 17.3 AND 17.4

MANUFACTURER'S DATA

AUXILIARY FUEL SYSTEM

AUXILIARY FUEL SYSTEM SERVICING

The auxiliary fuel tanks are interconnected with the basicfuel systemto allow gravity flow of auxiliary fuel into main fuel cells as fuel is consumed. The auxiliary fuel system is serviced simultaneously with the basic fuel systemthrough the singlefiller port locatedon the aft right side of the fuselage.

FUEL SYSTEM CAPACITIES

BASIC SYSTEM WITH LEFT OR RIGHT AUXILIARY TANK Total capacity: 419.1 U.S. gallons (1586.5 liters) Usablefuel: 412.1 U.S. gallons (1560.0 liters)

BASIC SYSTEM WITH BOTH AUXILIARY TANKS Total capacity: 500.8 U.S. gallons (1895.7 liters) Usable fuel: 493.8 U.S. gallons (1869.2 liters)

AUXILIARY FUEL LOADING TABLES

Fuel loading tables are presented for weight and balance computations in both English and Metric units. These tables shall be used in lieu of the tables for the basic fuel system when either or both auxiliary fuel tanks are installed. Weights and moments listed herein represent total fuel on board to include that contained in basic fuel cells. Refer to table 1-1 and 1-2 for English or 1-1M and 1-2M for Metric when both left and right auxiliary tanks are installed. Tables 1-3 and 1-4 for English or 1-3M and 14M for Metric apply to single auxiliary tank installed on left side, and tables 1-5 and 1-6 for English or 1-5M and 1-6M Metric, apply to single auxiliary tank installed on right side.

1-1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O Section 1


BHT-412-FMS-17.2. 17.3AND 17.4 Table 1-1.

Fuel Loading

With Left and Right

Auxiliary

Tanks -

Longitudinal

(English)

Longitudinal

Jet B or JP-4 (6.5 Lb/U.S.

Quantity (U.S. Gal)

Weight (Pounds)

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 207.9 210 220 230 240 241.0 250 260 270 280 290 300 310 320 330 340 350 360 370 374.5 380 390 400 410 420 426.4 430 440 450 460 470 480 490 493 8

65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1351 1365 1430 1495 1560 1567 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2405 2434 2470 2535 2600 2665 2730 2772 2795 2860 2925 2990 3055 3120 3185 3210

Gallon)

Jet A, A1 or JP5 (6.8 Lb/U.S.

CG (Inches)

Moment (In-Lb)

Quantity (U.S. Gal)

Weight (Pounds)

CG (Inches)

139.4 139.6 139.8 139.9 1399 1399 141.1 145.5 148.2 149.8 151.0 152.1 152.9 153.7 154.3 154.8 155.3 155.8 156 1 1564 156.8 157.0 156.5 1536 151 0 148 6 148.3 148.8 149.3 149.8 1502 150.7 151 1 151.5 151 9 152.2 152.5 152.8 153.1 153.3 153.4 152 9 151.9 151.0 1501 149.2 148.7 1488 149 1 149.4 149.7 1500 150.2 150.5 150.6

9061 18148 27261 36374 45468 53022 55029 66203 77064 87633 98150 108752 119262 129877 140413 150930 161512 172159 182637 193154 203840 212107 213623 219648 225745 231816 232386 241800 252317 262899 273364 284070 294645 305273 315952 326469 337025 347620 358254 368687 373376 377663 385067 392600 400017 407316 412196 415896 426426 436995 447603 458250 468624 479343 483426

10 20 30 40 50 '58.3 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 207.9 210 220 230 240 241.0 250 260 270 280 290 300 310 320 330 340 350 360 370 374 5 380 390 400 410 420 426.4 430 440 4J0 460 470 480 490 493.8

68 136 204 272 340 396 408 476 544 612 680 748 816 884 952 1020 1088 1156 1224 1292 1360 1414 1428 1496 1564 1632 1639 1700 1768 1836 1904 1972 2040 2108 2176 2244 2312 2380 2448 2516 2547 2584 2652 2720 2788 2856 2900 2924 2992 3060 3128 3196 3264 3332 3358

139.4 139 6 139 8 1399 139.9 139.9 141.1 145.5 148.2 149.8 151.0 152.1 152.9 1537 154.3 154.8 155 3 1558 156 1 156.4 1568 157.0 156.5 1536 151.0 1486 1483 148 8 149.3 149.8 1502 150.7 151.1 151 5 151 9 152 2 1525 152.8 153.1 153.3 153 4 1529 151 9 151.0 150 1 1492 148 7 148.8 149.1 149.4 149.7 150 0 150.2 1505 1506

Critical fuel amount for most forward cg condition. NOTE: All data above represents total fuel on board (basic and auxiliary),

Gallon)

Moment (In-Lb)

9479 18986 28519 38053 47566 55400 57569 69258 80621 91678 102680 113771 124766 135871 146894 157896 168966 180105 191066 202069 213248 221998 223482 229786 236164 242515 243064 252960 263962 275033 285981 297180 308244 319362 330534 341537 352580 363664 374789 385703 390710 395094 402839 410720 418479 426115 431230 435091 446107 457164 468262 479400 490253 501466 505715

based on nominal density at 15°C.


Section 1

BHT-412-FMS-17.2, 17.3 AND 17.4

Table 1-1M.

Fuel Loading

With Left and Right Auxiliary

Tanks -

Longitudinal

(Metric)

Longitudinal

Jet B or JP-4 (0.779

kg/liter)

Quantity (liters)

Weight (kg)

CG (mm)

40 80 120 160 200 220 7 240 280 320 360 400 440 480 520 560 600 640 680 720 760 787 0 800 840 880 912 3 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1400 1417 6 1440 1480 1520 1560 1600 1614.1 1640 1680 1720 1760 1800 1840 1869 2

31 2 62 3 93 5 124.6 1558 171 9 1869 218.1 249 3 280 4 311.6 342 7 373 9 405 1 436 2 467 4 498 5 529 6 560.8 592.0 613 0 623 1 654.2 685 4 710 8 716.7 747 7 778 9 8100 841 2 872 3 903 5 934 7 965.8 997 0 1028 1 1059.3 1090 4 1104 0 1121 6 1152.7 11839 1215 0 1246.0 1257 3 1277 3 1308 5 1339 7 1370 8 1402 0 1433 1 1455.9

3542 3547 3551 3552 3552 3553 3635 3727 3783 3825 3852 3877 3898 3915 3930 3944 3955 3965 3974 3982 3988 3965 3887 3817 3766 3770 3785 3797 3810 3822 3833 3844 3854 3864 3872 3880 3887 3895 3897 3882 3856 3832 3808 3784 3776 3781 3789 3797 3805 3813 3820 3825

Moment (kg-mm) 110510 220978 332019 442579 553402 610761 679382 812859 943102 1072530 1200283 1328648 1457462 1585967 1714266 1843426 1971568 2099864 2228619 2357344 2444644 2470592 2542875 2616172 2676873 2701959 2830045 2957483 3086100 3215066 3343526 3473054 3602334 3731851 3860384 3989028 4117499 4247108 4302288 4354051 4444811 4536705 4626720 4714864 4747565 4829471 4957907 5086841 5215894 5345826 5474442 5568818

Quantity (liters)

40 80 120 160 200 220.7 240 280 320 360 400 440 480 520 560 600 640 680 720 760 787 0 800 840 880 912 3 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1400 1417.6 1440 1480 1520 1560 1600 1614.1 1640 1680 1720 1760 1800 1840 1869 2

Jet A, A1 or JP-5 (0.815 kg/liter) Weight CG (kg) (mm) 32 6 3541 65 2 3547 97 8 3551 1304 3552 163.0 3552 1799 3553 1956 3635 228.2 3727 260 8 3783 293 3 3825 325.9 3852 358 5 3877 391.1 3898 423 7 3915 456 3 3930 488.9 3944 521 5 3955 554.1 3965 586.7 3974 619 3 3982 641 3 3988 651.9 3965 684.4 3887 717.0 3817 743.4 3766 749 8 3770 782.2 3785 814.8 3797 847 4 3810 880.0 3822 9129 3833 945.2 3844 977 8 3854 1010.4 3864 1043.0 3872 1075 6 3880 1108 1 3887 1140.7 3895 1155.1 3897 11734 3882 1205 9 3856 1238.5 3832 1271 1 3808 1303.7 3784 1315.2 3776 1336 3 3781 1368 9 3789 1401.5 3797 1434.1 3805 3813 1466 7 1499 3 3820 1523 3 3825

Critical fuel amount for most forward cg condition. NOTE: All data above represents total fuel on board (basic and auxiliary).

Moment (kg-mm) 115437 231264 347288 463181 578976 639185 711006 850501 986606 1121873 1255367 1389905 1524508 1658786 1793259 1928222 2062533 2197007 2331546 2466053 2557504 2584784 2660263 2736789 2799644 2826746 2960627 3093796 3228594 3363360 3499146 3633349 3768441 3904186 4038496 4173328 4307185 4443027 4501427 4555139 4649950 4745932 4840349 4933201 4966195 5052550 5186762 5321496 5456751 5592527 5727326 5826623


1-3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O Section

1


BHT-412-FMS-17.2. 17.3 AND 17.4 Table 1-2.

Fuel Loading With Left and Right Auxiliary Tanks - Lateral (English) Lateral


Quantity (U S. Gal.)

Weight (Pounds)

CG (Inches)

10 20 30 40 50 58 3 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 207.9 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 390 400 410 420 430 440 450 460 470 480 490 493.8

65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1351 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2405 2470 2535 2600 2665 2730 2795 2860 2925 2990 3055 3120 3185 3210

0 0 0 0 0 0 -0.04 0.06 -0 05 -0.04 -0 04 -0.03 -0 03 -0.03 -0.03 0.02 -0.02 -0.02 0.02 -0 02 -0 02 -0.02 -0 20 -0 33 0.42 -0.50 -0.50 0 48 -0.46 -0 45 0.44 -0 43 -0.41 -0.41 -0.40 -0 40 0 40 -0 39 -0 39 -0 36 0 35 -0 34 -0 33 0 31 -0.30 -0.30 -0 29 -0.29 -0 28 -0.27 -0.25 -0 25

Jet A, A1 or JP-5 (6.8 Lb/U.S.

Gallon)

Moment (In-Lb) 0 0 0 0 0 0 16 -27 -26 23 -26 21 -23 -25 27 -20 21 -22 -23 -25 26 -27 -27 472 -628 -780 -813 -811 -807 819 -829 -839 -826 -853 858 -884 -910 -913 938 -889 -887 -884 -879 -846 839 -858 -848 867 -855 -842 796 -803

Quantity U.S. Gall

Weight (Pounds)

CG (Inches)

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 207.9 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 390 400 410 420 430 440 450 460 470 480 490 493.8

68 136 204 272 340 397 408 476 544 612 680 748 816 884 962 1020 1088 1156 1224 1292 1360 1414 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2108 2176 2244 2312 2380 2448 2516 2584 2652 2720 2788 2856 2924 2992 3060 3128 3196 3264 3332 3358

0 0 0 0 0 0 -0.04 -0 06 -0 05 -0 04 -0.04 -0 03 -0 03 -0.03 0.03 -0 02 -0.02 -0.02 -0.02 -0.02 0.02 -0.02 -0.20 0 33 -0.42 -0.50 -0.50 0.48 -0.46 -0.45 -0 44 -0.43 -0 41 -0.41 -0.40 -0 40 -0.40 -0.39 0.39 -0.36 -0 35 0.34 -0.33 -0.31 -0 30 -0 30 -0.29 0.29 -0.28 -0.27 -0.25 -0.26

Critical fuel amount for most forward cg condition. NOTE: All data above represents total fuel on board (basic and auxiliary)

Gallon)

Moment (In-Lb) 0 0 0 0 0 0 -16 -29 -27 -25 -27 -22 -24 27 -29 -20 -22 -23 -24 26 -27 -28 286 -494 657 816 -850 -849 845 -857 -868 -877 -864 -292 -898 -925 952 -955 -981 930 -928 -925 920 -886 -877 -898 887 -907 -895 -881 -833 -840



Section 1

BHT-412-FMS-17.2. 17.3 AND 17.4

Table 1-2M.

Fuel Loading With Left and Right Auxiliary Tanks - Lateral (Metric)

Lateral

Jet B or JP-4 (0.779 kg/liter)

Quantity liters)

Weight (kg)

CG (mm)

40 80 120 160 200 220 7 240 280 320 360 400 440 480 520 560 600 640 680 720 760 787 0 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1400 1440 1480 1520 1560 1600 1640 1680 1720 1760 1800 1840 1869.2

31.2 62.3 93 5 124 6 155.8 171.9 186.9 218 1 249 3 280.4 311.6 342 7 373.9 405.1 436.2 467 3 498.5 529.6 560.8 592.0 613.0 623.1 654.2 685.4 716.7 7477 778.9 810.0 841 2 872.3 9035 934 7 965 8 997.0 1028.1 1059.3 1090 4 1121.6 1152.7 1183.9 1215.0 1246.0 1277 3 1308.5 1339.7 1370.8 1402.0 1433.1 1455.9

0 0 0 0 0 0 2 2 1 -1 -1 -1 -1 1 -1 1 1 -1 1 -1 1 -5 9 -11 -13 -13 -12 12 12 11 -10 -10 -10 -10 10 10 10 9 -9 -8 -8 8 -8 8 -7 -7 -7 -6 6

Jet A, A1 or JP-5 (0.815

Moment (kg-mm)

0 0 0 0 0 0 -374 436 249 280 312 -343 -374 -405 -436 467 -499 530 -561 -592 -613 3116 5888 -7539 9317 -9720 -9347 -9720 -10094 9595 -9035 9347 -9658 -9970 -10281 -10593 10904 -10094 10374 -9471 9720 -9968 -10218 10468 -9378 9596 -9814 -8599 -8735

Quantity liters)

Weight (kg)

CG (mm)

40 80 120 160 200 220 7 240 280 320 360 400 440 480 520 560 600 640 680 720 760 787.0 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1400 1440 1480 1520 1560 1600 1640 1680 1720 1760 1800 1840 1869.2

32 6 65.2 97.8 130.4 163 0 179.9 195.6 228.2 260.8 293 3 325.9 358.5 391.1 423.7 456.3 488 9 521.5 554.1 586 7 619.3 641.3 -651.9 684 4 7170 749.8 7822 814.8 847.4 8800 912.9 9452 977 8 1010.4 1043.0 1075.6 1108.1 1140 7 11734 1205.9 1238.5 1271.1 1303.7 1336 3 1368.9 1401.5 1434.1 1466 7 1499.3 1523.2

0 0 0 0 0 0 -2 2 1 -1 1 -1 -1 -1 -1 1 -1 1 -1 -1 -1 5 9 11 13 -13 -12 -12 12 11 10 10 -10 -10 -10 10 10 9 9 -8 -8 -8 8 8 7 7 -7 -6 -6

Critical fuel amount for most forward cg condition. NOTE: All data above represents total fuel on board (basic and auxiliary).

kg/liter)

Moment (kg-mm)

0 0 0 0 0 0 391 -456 -261 293 -326 359 391 424 456 -489 522 -554 587 -619 641 -3260 -6160 -7887 -9747 10169 -9778 10169 -10560 10042 9452 -9778 10104 -10430 10756 -11081 -11407 -10561 -10853 9908 10169 10430 -10690 -10951 -9811 10039 10267 -8996 -9139


1-5

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section


BHT-412-FMS-17.2, 17.3 AND 17.4 Table 1-3.

Fuel Loading

With

Left Auxiliary

Tank -

Longitudinal

(English)

Longitudinal


Jet A, Al

Gallon)

or JP-5 (6.8 Lb/U.S.

Gallon)

Quantity (U.S.Gal)

Weight (Pounds)

CG (Inches)

Moment (In-Lb)

Quantity (U.S. Gal)

Weight (Pounds)

CG (Inches)

Moment (In-Lb)

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 160 170 173.9 180 190 200 207.1 210 220 230 240 250 260 270 280 290 300 308.8 310 320 330 340 350 360 360.7 370 380 390 400 410 412.1

65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1130 1170 1235 1300 1346 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2007 2015 2080 2145 2210 2275 2340 2345 2405 2470 2535 2600 2665 2679

139.4 139.6 139.8 1399 139.9 139.9 141.1 1458 148.7 150.7 152.3 153.6 154.7 155.7 156.5 157.2 157.8 158.4 158.6 156.2 153.1 150.1 148.2 148.4 149.2 1500 150.7 151.4 152.0 152 6 153.1 153.6 154.0 154 4 154.2 153.0 151.8 150.7 1497 1486 148.6 1490 149.6 150.1 150.5 160.8 150.9

9061 18148 27261 36374 45468 53022 55029 66339 77324 88160 98995 109824 120666 131567 142415 153270 164112 175032 179218 182754 189079 195130 199477 202566 213356 224250 235092 246025 256880 267813 278642 289536 300300 309881 310713 318240 325611 333047 340568 347724 348467 358345 369512 380504 391300 401882 404261

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 160 170 173 9 180 190 200 207.1 210 220 230 240 250 260 270 280 290 300 308.8 310 320 330 340 350 360 360.7 370 380 390 400 410 412.1

68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1088 1156 1183 1224 1292 1360 1408 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2100 2108 2176 2244 2312 2380 2448 2453 2516 2584 2652 2720 2788 2802

139.4 139.6 139.8 139.9 139.9 139.9 141.1 145.8 1487 150.7 152.3 153.6 154.7 155.7 156.5 157.2 157.8 1584 158.6 156 2 153 1 150.1 148.2 1484 149 2 150.0 150 7 151.4 1520 152.6 153.1 1536 154 0 154 4 1542 1530 151.8 150.7 149.7 148.6 148.6 1490 1496 150 1 150.5 150.8 150.9

9479 18986 28519 38053 47566 55540 57569 69401 80893 92228 103564 114893 126235 137639 148988 160344 171686 183110 187624 191189 197805 204136 208666 211915 223203 234600 245942 257380 268736 280174 291502 302899 314160 324240 325054 332928 340639 348418 356286 363773 364516 374884 386566 398065 409360 420430 422822

'Critical NOTE:

fuel amount for most forward cg condition. All data above represents total fuel on board (basic and auxiliary).

based on nominal density at 15 C.



BHT-412-FMS-17.2, 17.3 AND 17.4

Table 1-3M.

Fuel Loading

With

Left Auxiliary

(Metric)

Longitudinal

Tank -

Longitudinal

Jet B or JP-4

Quantity (liters)

40 80 120 160 200 220 7 240 280 320 360 400 440 480 520 560 600 640 658.3 680 720 760 784 0 800 840 880 920 960 1000 1040 1080 1120 1160 1168 9 1200 1240 1280 1320 1360 1365.4 1400 1440 1480 1520 1560 0

Critical NOTE:

Weight (kg)

31.2 62 3 93 5 124.6 155.8 171.9 186 9 218.1 249.3 280 4 311 6 342.7 373.9 405.0 436.2 467 3 498 5 512 5 529.6 560.8 592.0 610.4 623 1 654.2 685.4 716.7 747.7 778 9 810.0 841 2 872 3 903.5 910.2 934.7 965 8 997.0 1028.1 1059 3 1063.5 1090.4 1121.6 1152 7 1183 9 1215.0

(0.779

Jet A, Al

kg/liter)

CG (mm)

Moment (kg-mm)

3542 3547 3551 3552 3552 3553 3635 3739 3800 3848 3889 3923 3948 3971 4001 4006 4021 4027 3973 3886 3809 3763 3772 3793 3815 3834 3852 3868 3882 3896 3909 3919 3922 3895 3863 3833 3805 3778 3775 3787 3800 3812 3824 3834

110510 220978 332019 442579 553402 610761 679382 815476 947340 1078979 1211812 1344412 1476157 1608255 1745236 1872004 2004469 2063838 2104101 2179269 2254928 2296935 2350333 2481381 2614801 2747828 2880140 3012785 3144420 3277315 3409821 3540817 3569804 3640657 3730885 3821501 3911921 4002035 4014713 4129345 4262080 4394092 4527234 4658310

Quantity (liters)

40 80 120 160 200 220 7 240 280 320 360 400 440 480 520 560 600 640 658.3 680 720 760 784 0 800 840 880 920 960 1000 1040 1080 1120 1160 1168.9 1200 1240 1280 1320 1360 1365 4 1400 1440 1480 1520 1560 0

fuel amount for most forward cg condition. All data above represents total fuel on board (basic and auxiliary),

or JP-5 (0.815

Weight (kg)

32 6 65 2 97 8 130.4 163.0 179.9 195 6 228 2 260 8 293.3 325 9 358.5 391.1 423.7 456 3 488.9 521.5 536 5 554.1 586 7 619.3 638.6 651.9 684.4 717.0 749 8 782.2 814.8 847.4 880 0 912 9 945.2 952 5 977.8 1010 4 1043 0 1075.6 1108 1 1112 6 1140.7 1173.4 1205.9 1238 5 1271 0

kg/liter)

CG (mm)

Moment (kg-mm)

3541 3547 3551 3552 3552 3553 3635 3739 3800 3848 3889 3923 3948 3971 4001 4006 4021 4027 3973 3886 3809 3763 3772 3793 3815 3834 3852 3868 3882 3896 3909 3919 3922 3895 3863 3833 3805 3778 3775 3787 3800 3812 3824 3834

115437 231264 347288 463181 578976 639185 711006 853240 991040 1128618 1267425 1406396 1544063 1682513 1825656 1958533 2096952 2160486 2201439 2279916 2358914 2403052 2458967 2595929 2735355 2874733 3013034 3151646 3289607 3428480 3568526 3704239 3735705 3808531 3903175 3997819 4092658 4186402 4200065 4319831 4458920 4596891 4736024 4873014




BHT-412-FMS-17.2, 17.3 AND 17.4 Table 1-4.

Fuel Loading

with

Left Auxiliary

Tank -

Lateral (English)

Lateral


Quantity (U.S. Gal) 10 20 30 40 50 58 3 60 70 80 90 100 110 120 130 140 150 160 170 175.4 180 190 200 2071 210 220 230 240 250 260 270 280 290 300 308.8 310 320 330 340 350 360 360 7 370 380 390 400 410 412.1

' Critical

NOTE.

1-8

Weight (Pounds) 65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1140 1170 1235 1300 1346 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2007 2015 2080 2145 2210 2275 2340 2345 2405 2470 2535 2600 2665 2679

CG (Inches) 0 0 0 0 0 0 -0 03 -0.45 -1 45 2 25 2 97 -3.55 4 04 -4 45 4 80 -5.10 -5 34 -5 55 579 5.70 -5 58 -5 40 -5 32 -5.38 -5 55 5 70 -5 85 -5.97 -6.10 6 20 630 -6 40 6 47 -6.54 6 50 -6.30 -610 -5 95 5 80 -5.60 5 60 -5 68 -5 75 -5 85 -5 94 -6 03 6 03

fuel amount for most forward

Gallon)

Moment (In-Lb)

0 0 0 0 0 0 12 205 -754 -1316 1931 2538 -3151 3760 -4368 4973 5554 6133 -6601 -6669 -6891 7020 7161 -7344 -7937 8522 9126 9701 -10309 -10881 -11466 -12064 -12617 13126 13098 -13104 -13085 13150 13195 -13104 -13132 13660 14203 -14830 -15444 16070 16154

Jet A, Al or JP-5 (6.8 Lb/U.S.

Quantity (U.S. Gal) 10 20 30 40 50 58 3 60 70 80 90 100 110 120 130 140 150 160 170 175.4 180 190 200 207 1 210 220 230 240 250 260 270 280 290 300 308.8 310 320 330 340 350 360 360 7 370 380 390 400 410 4121

Weight (Pounds) 68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1088 1156 1193 1224 1292 1360 1408 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2100 2108 2176 2244 2312 2380 2448 2453 2516 2584 2652 2720 2788 2802

CG (Inches) 0 0 0 0 0 0 0.03 -0 45 1 45 2 25 2.97 3 55 -4.04 4 45 4.80 -5 10 -5 34 5 55 5.79 -5 70 -5 58 5.40 5.32 -5 38 -5 55 5.70 -5.85 -5 97 -6 10 6.20 630 6 40 6 47 6.54 6 50 -6 30 6 10 -5.95 5 80 5 60 5 60 5.68 5 75 5 85 5 94 6.03 6.03

Gallon)

Moment (In-Lb)

0 0 0 0 0 0 -12 -214 -789 1377 -2020 -2655 -3297 3934 -4570 -5202 5810 -6416 -6907 -6977 -7209 -7344 -7491 -7683 8303 -8915 -9547 10149 -10785 -11383 -11995 12621 -13199 -13734 -13702 -13709 13688 -13756 -13804 -13709 -13737 -14291 -14858 15514 16157 -16812 -16896

cg condition.

All data above represents total fuel on board (basic and auxiliary),



Section 1

BHT-412-FMS-17.2, 17.3AND 17.4

Table 1-4M.

Fuel Loading

with Left Auxiliary

Tank -

Lateral (Metric)

Lateral

Jet A, A1 or JP-5 (0.815 kg/liter)

Jet B or JP-4 (0 779 kg/liter) Weight (kg)

Quantity (liters) 40 80 120 160 200 220 240 280 320 360 400 440 480 520 560 600 640 664 680 720 760 784 800 840 880 920 960 1000 1040 1080 1120 1160 1168 1200 1240 1280 1320 1360 1365 1400 1440 1480 1520 1560

Critical

NOTE.

7

0

0

9

4

31 2 62 3 93 5 124 6 1558 171 9 186 9 218 1 249 3 280 4 311 6 342 7 373 9 4050 436 2 467 3 498.5 517.0 529 6 560 8 592 0 6104 623.1 654 2 685 4 7167 747 7 778 9 8100 841 2 872 3 903 5 9102 934 7 965 8 997 0 1028 1 1059 3 1063 5 1090 4 1121 6 1152 7 1183 9 12150

fuel amount

CG (mm)

Moment (kg-mm)

0 0 0 0 0 0 -1 21 -35 -67 -84 99 -110 -119 -128 -135 -140 -147 145 -142 -138 -135 137 -141 -145 150 -152 155 -158 161 163 166 -166 162 -157 -152 -147 -142 142 144 146 -149 -151 -153

0 0 0 0 0 0 187 -4580 -8726 -18787 26174 -33927 -41129 -48195 55834 63086 -69790 75999 -76792 79634 -81696 82404 -85365 -92242 -99383 -107505 113650 -120730 127980 -135433 -142185 149981 151093 -151421 -151631 151544 -151131 -150421 -151017 157018 163754 -171752 -178769 185895

formost forward

Quantity (liters)

40 80 120 160 200 220 7 240 280 320 360 400 440 480 520 560 600 640 664 0 680 720 760 784 0 800 840 880 920 960 1000 1040 1080 1120 1160 11689 1200 1240 1280 1320 1360 1365 4 1400 1440 1480 1520 1560

Weight (kg)

32 6 65 2 978 1304 163.0 179 9 195 6 228 2 260 8 293 3 325 9 358 5 391 1 423 7 456 3 488 9 521 5 541 0 554 1 586 7 619 3 638 6 651 9 684 4 7170 749 8 782 2 814 8 8474 880.0 9129 945 2 952 5 9778 10104 10430 1075 6 1108 1 1112 6 1140 7 1173 4 1205 9 1238 5 1271 0

CG (mm)

0 0 0 0 0 0 -1 -21 35 -67 84 -99 -110 -119 128 135 -140 147 -145 -142 -138 135 -137 -141 -145 -150 152 -155 158 -161 -163 -166 -166 162 157 152 147 -142 -142 144 -146 -149 -151 153

Moment (kg-mm)

0 0 0 0 0 0 -196 -4792 -9128 19651 27376 -35492 -43021 -50420 58406 -66002 73010 79527 -80345 -83311 -85463 -86211 -89310 -96500 -103965 -112470 -118894 126294 133889 -141680 148803 -156903 -158115 -158404 158633 158536 -158113 157350 -157989 164261 -171316 179679 -187014 -194463

cg condition

All data above represents total fuel on board (basic and auxiliary).


1-9


Section 1

BHT-412-FMS-17.2, 17.3 AND 17.4

Table 1-5.

Fuel Loading with Right Auxiliary Tank - Longitudinal (English) Longitudinal


Quantity (U S. Gal)

Weight (Pounds)

Jet A, A1 or JP-5 (6.8 Lb/U.S

Gallon)

CG (Inches)

Moment (In-Lb)

Quantity (U.S Gal)

139.4 139 6 139 8 139.9 139 9 139 9 141 1 145 8 148 7 150 7 152 3 1536 154 7 155 7 156 5 157.2 1578 1584 1586 156 2 153 1 150 1 148 2 148.4 149 2 150 0 150 7 151 4 152 0 152 6 1531 1536 154 0 154 4 154 2

9061 18148 27261 36374 45468 53022 55029 66339 77324 88160 98995 109824 120666 131567 142415 153270 164112 175032 179218 182754 189079 195130 199477 202566 213356 224250 235092 246025 256880 267813 278642 289536 300300 309881 310713

10 20 30 40 50 '58 3 60 70 80 90 100 110 120 130 140 150 160 170 1739 180 190 200 2071 210 220 230 240 250 260 270 280 290 300 308 8 310

Gallon)

CG (Inches)

Moment (In-Lb)

68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1088 1156 1183 1224 1292 1360 1408 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2100 2108

139 4 139 6 139 8 139.9 139 9 139 9 141 1 145 8 148 7 150.7 152 3 1536 154 7 155 7 156 5 157 2 1578 1584 1586 156 2 153 1 150 1 148 2 1484 149 2 150 0 150 7 151 4 152 0 152 6 153 1 1536 154 0 154 4 154 2

9479 18986 28519 38053 47566 55540 57569 69401 80893 92228 103564 114893 126235 137639 148988 160344 171686 183110 187624 191189 197805 204136 208666 211915 223203 234600 245942 257380 268736 280174 291502 302899 314160 324240 325054

Weight (Pounds)

10 20 30 40 50 '58 3 60 70 80 90 100 110 120 130 140 150 160 170 1739 180 190 200 2071 210 220 230 240 250 260 270 280 290 300 308 8 310

65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1130 1170 1235 1300 1346 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2007 2015

320

2080

153.0

318240

320

2176

153 0

332928

330 340 350 360 360 7 370 380

2145 2210 2275 2340 2345 2405 2470

151 8 150 7 149 7 148 6 148 6 149.0 149 6

325611 333047 340568 347724 348467 358345 369512

330 340 350 360 360 7 370 380

2244 2312 2380 2448 2453 2516 2584

151 150 149 148 148 149 149

340639 348418 356286 363773 364516 374884 386566

390

2535

150 1

380504

390

2652

150 1

398065

400 410 412 1

2600 2665 2679

150 5 150 8 150 9

391300 401882 404261

400 410 412 1

2720 2788 2802

150 5 150 8 150 9

409360 420430 422822

8 7 7 6 6 0 6

* Critical fuel amount for most forward cg condition. NOTE:

1-10


based on nominal density at 15' C


Section 1

BHT-412-FMS-17.2, 17.3 AND 17.4

Table 1-5M.

Fuel Loading

with

Right Auxiliary

Tank -

Longitudinal

(Metric)

Longitudinal

Jet B or JP-4 (0.779

Quantity (liters)

Weight (kg)

40 80 120 160 200 220.7 240 280 320 360 400 440 480 520 560 600 640 658 3 680 720 760 784 0 800 840 880 920 960 1000 1040 1080 1120 1160 11689 1200 1240 1280 1320 1360 1365 4 1400 1440 1480 1520 1560.0

Critical

NOTE:

31 2 62.3 93 5 124 6 155.8 171 9 186 9 218.1 249 3 280.4 311.6 342.7 373.9 405.0 436 2 467 3 498 5 5125 529 6 560 8 592.0 610.4 623.1 654 2 685 4 716.7 747 7 778 9 8100 841.2 8723 903.5 910.2 934.7 965.8 997.0 1028.1 1059 3 1063.5 10904 1121.6 1152.7 11839 12150

kg/liter)

CG (mm)

3542 3547 3551 3552 3552 3553 3635 3739 3800 3848 3889 3923 3948 3971 4001 4006 4021 4027 3973 3886 3809 3763 3772 3793 3815 3834 3852 3868 3882 3896 3909 3919 3922 3895 3863 3833 3805 3778 3775 3787 3800 3812 3824 3834


Moment (kg-mm)

110510 220978 332019 442579 553402 610761 679382 815476 947340 1078979 1211812 1344412 1476157 1608255 1745236 1872004 2004469 2063838 2104101 2179269 2254928 2296935 2350333 2481381 2614801 2747828 2880140 3012785 3144420 3277315 3409821 3540817 3569804 3640657 3730885 3821501 3911921 4002035 4014713 4129345 4262080 4394092 4527234 4658310

Quantity (liters)

40 80 120 160 200 220.7 240 280 320 360 400 440 480 520 560 600 640 658 3 680 720 760 784.0 800 840 880 920 960 1000 1040 1080 1120 1160 1168.9 1200 1240 1280 1320 1360 1365.4 1400 1440 1480 1520 1560 0

Weight (kg)

CG (mm)

Moment (kg-mm)

32.6 65 2 97.8 130.4 163 0 179 9 195 6 228 2 260.8 293.3 325.9 358.5 391.1 4377 456.3 488.9 521 5 536.5 554.1 586.7 6193 638.6 651.9 684.4 717.0 749.8 782.2 814.8 847 4 880.0 9129 945.2 9525 977.8 1010.4 1043.0 1075.6 1108.1 1112.6 1140.7 1173.4 1205.9 1238.5 1271.0

3541 3547 3551 3552 3552 3553 3635 3739 3800 3848 3889 3923 3948 3971 4001 4006 4021 4027 3973 3886 3809 3763 3772 3793 3815 3834 3852 3868 3882 3896 3909 3919 3922 3895 3863 3833 3805 3778 3775 3787 3800 3812 3824 3834

115437 231264 347288 463181 578976 639185 711006 853240 991040 1128618 1267425 1406396 1544063 1682513 1825656 1958533 2096952 2160486 2201439 2279916 2358914 2403052 2458967 2595929 2735355 2874733 3013034 3151646 3289607 3428480 3568526 3704239 3735705 3808531 3903175 3997819 4092658 4186402 4200065 4319831 4458920 4596891 4736024 4873014

fuel amount for most forward cg condition.


0


1-11


Section

BHT-412-FMS-17.2,

17.3 AND 17.4

Table 1-6.

Fuel Loading with Right Auxiliary Tank -Lateral

(English)

Lateral


Quantity (U.S. Gal) 10 20 30 40 50 58 3 60 70 80 90 100 110 120 130 140 150 160 170 173.9 180 190 200 207.1 210 220 230 240 250 260 270 280 290 300 308.8 310 320 330 340 350 360 360.7 370 380 390 400 410 412.1

* Critical NOTE:

Weight (Pounds) 65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1040 1105 1130 1170 1235 1300 1346 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2007 2015 2080 2145 2210 2275 2340 2345 2405 2470 2535 2600 2665 2679

CG (Inches)

0 0 0 0 0 0 0.03 0.34 1 30 2.22 2.90 3.48 3 95 4 43 4.71 5.00 5.28 5.50 5.60 5.15 4.73 4.36 4.12 4 20 4.45 4.63 481 5.00 5.13 5.28 5.40 5.60 5.65 5.74 5 72 5 53 5.36 5.20 5.05 4.95 4.92 501 5.10 5.20 5.30 541 5.43

fuel amount for most forward All data above represents

Gallon)

Jet A, A1 or JP-5 (6.8 Lb/U.S.

Moment (In-Lb)

0 0 0 0 0 0 -12 155 676 1299 1885 2488 3081 3743 4286 4875 5491 6078 6328 6026 5842 5668 5546 5733 6364 6922 7504 8125 8670 9266 9828 10556 11018 11520 11526 11502 11497 11492 11489 11583 11537 12049 12597 13182 13780 14418 14547

Quantity (U.S. Gal) 10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 160 170 173.9 180 190 200 207.1 210 220 230 240 250 260 270 280 290 300 308.8 310 320 330 340 350 360 360.7 370 380 390 400 410 412.1

Weight (Pounds) 68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1088 1156 1183 1224 1292 1360 1408 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2100 2108 2176 2244 2312 2380 2448 2453 2516 2584 2652 2720 2788 2802

CG (Inches) 0 0 0 0 0 0 -0.03 0.34 1.30 2.22 2.90 3.48 3.95 4.43 4.71 5.00 5.28 5.50 5.60 5.15 4.73 4.36 412 4.20 4.45 4.63 4.81 5.00 5.13 5.28 5.40 5.60 5.65 5.74 5.72 5.53 5.36 5.20 5.05 4.95 4.92 5.01 5.10 5.20 5.30 5.41 5.43

Gallon)

Moment (In-Lb)

0 0 0 0 0 0 -12 162 707 1359 1972 2603 3223 3916 4484 5100 5745 6358 6625 6304 6111 5930 5801 5998 6657 7241 7850 8500 9070 9694 10282 11043 11526 12054 12058 12033 12028 12022 12019 12118 12069 12605 13178 13790 14416 15083 15215

cg condition.

total fuel on board (basic and auxiliary),


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O . 412 ROTORCRAFT MANUFACTURER'S DATA

Section 1

BHT-412-FMS-17.2,17.3 AND 17.4

Table 1-6M.

Fuel Loading with Right Auxiliary Tank -

Lateral (Metric)

Lateral


Quantity

Weight

CG


Moment

Quantity

Weight

CG

Moment

Criticalfuel amountfor mostforward cg condition.

NOTE: All data above represents total fuel on board (basic and auxiliary), based on nominal density at 15°C.

1-13

Section 1


BHT-412-FMS-17.2,17.3 AND 17.4

LOADING EXAMPLES

The CG examples shown below apply only to helicopters with standard seating which are equipped with both tanks of the 412-706-007 Auxiliary Fuel Kit. The loadings are based on standard 170 pound people for both crew and passengers. At empty weights below 7400 pounds there are no restrictions on passenger or fuel loadings except as imposed by the gross weight CG limits of the helicopter.

When flying with 2 crew members and 9 passengers add weight in the baggage compartment as required to maintain C.G. within the forward limits.

Examples:

With a ballasted weight empty of 7500 pounds. 20 pounds is required in the baggage compartment. With a ballasted weight empty of 7700 pounds, 40 pounds is required in the baggage compartment.

With helicopter weights at or above 7400 pounds the following procedure should be applied: Ballast the helicopter to the most aft weight

1-14

WARNING

BAGGAGE BAG

COMPARTMENT

WEIGHT MUST BE REMOVED FOR SINGLE PILOT OPERATION.

T N I R P N O R I ROTORCRAFT O T F C FLIGHT MANUAL U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-18.2 AND 18.3

MODEL

33108 -33213 36001 - 36019

AND

33214 - 33999

36020 AND SUB

SUPPLEMENTFOR LOUDHAILEROPERATIONS 412-899-143

CERTIFIED NOVEMBER17, 1983

This supplement shall be attached to the Model 412 Flight Manual (BHT-412-FM-2 or -3) when the 412-899-143 Loudhailer has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations. Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopteri ii ;lI A Subsdiary

COPYRIGHT NOTICE

POST OFFICE

BOX

*42 .

or Textron Inc

FORT WORTH TEXAS 76101

COPYRIGHT 1991

BELL HELICOPTER TEXTRON INC

ANDBELLHELICOPTERTEXTRON,

ADIVISION OFTEXTRON CANADA LTD ALLRIGHTS RESERVED

17

NOVEMBER

REISSUED

-

1983

8 OCTOBER 1991



These data are proprietary to Bell Helicopter Textron, Inc. Disclosure. reproduction, or use of these data for any purpose other than helicopter operation and/or maintenance is forbidden without prior written authorization from Bell Helicopter Textron, Inc.

Additional copies of this publication may be obtained by contacting:

CommercialPublicationDistributionCenter Bell HelicopterTextron P.O.Box482 Fort Worth, Texas76101

PN


BHT-412-FMS-18.2 AND 18.3

LOG OF REVISIONS

Original ........ Reissued ....... Reissued .......

0 ..... November 17, 1983 0 .......... May 8. 1989 0 ...... October8, 1991

LOG OF PAGES

REVISION NO.

PAGE

Title .......................... PN ........................... A/B ........................... i/ii ............................ 1-3/4 .......................

PAGE

REVISION NO.

0 0 0 0 0

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT.WORTH. TEXAS 76193-0170

NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of supersededpages.

A/B


BHT-412-FMS-18.2 AND 18.3

INTRODUCTION

crew to direct ground personnel while remaining airborne. The kit contains a speaker assembly, amplifier, switches, and the necessary hardware to complete the installation. Use of Loudhailer is controlled through pilot or copilot control panel. Optional configurations allow use of a remote (hand held) microphone and/or a tape recorder.

i/ii



Actual weight change shall be determined after Loudhailer is installed and ballast readjusted if necessary to return empty weight CG within allowable limits.


IFR operation is prohibited with Loudhailer installed.

BHT-412-FMS-18.2 AND 18.3


BHT-412-FMS-18.2 AND 18.3

BEFOREEXTERIORCHECK

PA circuit breakers - Check in.

PA SYSTEM PWR switch - OFF.

SIREN/MOM switch - OFF.

EXTERIORCHECK

TRILL/MOM switch - OFF.

PASYSTEMGAINcontrolswitch - OFF.

FUSELAGE

Fuselage underside - Check security and wiring connections of Loudhailer. LOUDHAILER OPERATION

WARNING

USE EXTREME CARE DURING OF OPERATION GROUND LOUDHAILER TO PREVENT INJURY TO PERSONNEL.GROUND SUPPORT PERSONNELIN VICINITY OF HELICOPTER SHOULD WEAR PROTECTIVEHEARINGDEVICES.

2

NOTE

OFF position is the minimum gain preset at the remote amplifier located In the baggage compartment.

PASYSTEMPWRswitch - PWR.

Rotate control switch on communications controlpanelto HAIL/AUX. PAmodeselect- As desired.

PA SYSTEM PWR switch - OFF, when Loudhaileroperationis completed.

T N I R Section3 P N O R I O CT F U T O OD N R E Section 4 P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-18.2 AND 18.3



3/4

T N I R P N ROTORCRAFT O R I FLIGHTMANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-19.2, 19.3 AND 19.4

MODELS

33108 -33213 36001 -36019

AND

36020-

36086

AND 36087 AND SUB SUPPLEMENTFOR SOFT INTERIOR 412-705-510 CERTIFIED 28 MARCH 1985

This supplement shall be attached to the Model 412 Flight Manual (BHT-412-FM-2, or -3) or Model 412EP Flight Manual (BHT-412-FM-4) when the 412-705-510 Soft interior has been installed.

The information contained hereinsupplementsthe information of the basic Flight Manual. For Limitations, Procedures, and PerformanceData not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter COPYRIGHT

COPYRIGHT COPYRIGHT 1994

NOTICE

BELL HELICOPTER TEXTRON INC ANDBELLHELICOPTER TEXTRON A DIVISIONOFTEXTRON CANADA LTD ALLRIGHTSRESERVED

NOTICE

Subsidiaryof TextronInc

POST OFFICE BOX 482 . FORT WORTH. TEXAS 76101

REISSUE -

5 OCTOBER1994

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-19.2, 19.3 AND 19.4

Additional copies of this publication may be obtained by contacting: CommercialPublication DistributionCenter Bell HelicopterTextron P.O. Box 482 Fort Worth, Texas76101

PN


BHT-412-FMS-19.2, 19.3 AND 19.4

LOG OF REVISIONS

Original .......... Revision ........ Reissue .......... Reissue ..........

0 1 0 0

........ ...... ... ........ ........

28 Mar 85 08 May 89 08 Oct 91 05 Oct 94

LOG OF PAGES

REVISION NO.

PAGE

Title .......................... PN ........................... A/B ........................... i/ii ............................ 1/2 ...........................

PAGE

REVISION NO.

0 0 0 0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT. WORTH, TEXAS 76193-0170

NOTE

Revisedtext is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

A/B

FAA APPROVED


INTRODUCTION

The soft interior, when installed, will permit the helicopter to be flown with doors off/open with an airspeed limitation of 100 KIAS.

i/ii


FAA APPROVED

Section

1


Actual weight change shall be determined after Soft Interior is installed and ballast readjusted if necessary to return empty weight CG within allowablelimits.

In all cases, door configuration

shall be

symmetrical for both sides of the fuselage. NOTE

Opening or removing doors shifts

helicopter center of gravity and reduces

DOORS OPEN OR REMOVED

Remove both seat backs from outboard facing seatson each side of transmissionpylon.

Flight operation is approved for the following alternative configurations during VFR conditions only:

VNE. Refer to Weight

and Balance

section in Manufacturer's Data and to Airspeed Limitations.

AIRSPEED LIMITATIONS

VNEwith doors open or removed is 100 KIAS.

Both crew doors removed.

Both sliding doors locked open or removedwith both hingedpanels installed or removed.

1/2



BHT-412-FMS-21



LOGOF REVISIONS

LOG OF PAGES

Page

Revision No.

Page

Revision No.

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATION DIRECTORATE AIRCRAFTCERTIFICATION SERVICE DEPARTMENTOF TRANSPORTATION SOUTHWESTREGION,FORTWORTH, TEXAS

NOTE: Revisedtext is indicated by a black vertical line. Insert latest revision pages;dispose of supersededpages.

A

Use or disclosure of data contained on this page is subject to

the restrictionon the title pageof this document.



Introduction

BHT-412-FMS-21

INTRODUCTION

The GNS-500A/S3 is a very low frequency VLF/OMEGA radio navigation system that provides great circle point-to-point navigation on a worldwide basis. It also displays to the flight crew an assessment of its ability to navigate and will automatically revert to dead reckoning during periodsof inadequatesignalreception.

The GNS-500A/S3 consists of a control display unit (CDU) mounted on the pedestal, a computer receiver unit (CRU) and an optional equipment unit (OEU) located in the right avionics compartment. A NAV-1/VLFswitch is located on the pilot instrument panel next to the HSI. It is a combination 2 position pushbutton and annunciator. Pressing the pushbutton will couple the HSI to the NAV-1 receiver or GNS-500A/S3. The NAV-1 or VLF annunciator light will illuminate to identify the coupled mode. BRG PTR select switches are located on both the pilot and copilot instrument panels and allows selection of VLF or ADF bearing to be displayed on the pilot or copilot HSI.

Useor disclosure of data containedon this pageissubjectto the restrictionon the title pageof this document.



Section 1

BHT-412-FMS-21

OPERATIONALLIMITATIONS

The following GNS Operators Manual must be immediately availableto the flight crew whenever navigation is predicated on the use of the GNS-500A/S3: Operator's manual Report 1080 dated 9-1-80 for the -3A Program and CRT/CDU.

During RNAV operation of the GNS-500A/S3, additional navigation equipment required for the specific type of operation must be installed and operable.

The GNS-500A/S3 position information must be checked for accuracy (reasonableness) prior to use as a means of navigation and under the following conditions.

Operator's Manual Report 1080 dated 1-19-81 for the -3B Programand CRT/CDU.

Priorto each compulsory reporting point during IFRoperation when not under radar surveillance or control.

Provided the GNS-500A VLF/Omega navigation system is receiving usable signals from at least two Omega navigation stations, it is approved for:

At or prior to arrival at each enroute waypoint during IFR operation along approved RNAV routes.

VLF/IFR RNAV operation within the common boundaries of the United States and Alaska in accordance with the EN ROUTEcriteria of AC 90-45A or the criteriaof AC 20-101 B.

Operation as a meansto update self-contained navigation systems, such as INS or Doppler,in accordance with AC 120-31A in the areas between Latitudes 85°N to 55°S, with the exception of the area above 45°N Latitude bounded by Longitudes 30°E and 120°N extending across the Asian Continent. Operation as sole means of long range navigation in accordance with AC 120-37 in the areas between Latitudes 85°N to 55°S, with the exception of the area above 45°N Latitude bounded by Longitudes 30°E and 120°E extending across the Asian Continent.

Prior to requesting off-airway routing, and at hourly intervals thereafter during RNAV operationoff of approved RNAVroutes.

During period of Dead Reckoning,navigation shall not be predicatedon the use of the GNS-500A/S3 for RNAVoperation.

Following a period of Dead Reckoning, the helicopter position should be verified by visually sighting ground reference points and/or by using other navigation equipment such as VOR, DME, Tacan,NDB. or radar fix.

The GNS-500A/S3 may not be used for navigation in terminal areas or during departures from, or approaches,to, airports.

Useor disclosureof data containedonthis pageis subjectto therestrictionon thetitle pageof this document.

1

T N I R P N O R I O T F C U T OSectionO2 D N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHTMANUAL

Sections 1 & 2


BHT-412-FMS-21

LIMITATIONS(Cont) OPERATIONAL

Enroute navigation shall not be predicated on the GNS-500A/S3 during the period that DR is illuminated.

CHECK PREFLIGHT

NOTE

EXTERIORCHECK

When the HSI is coupled to the VLF.the courseset knob is disabled.

ENGINEAREA, AFT COMPARTMENTS AND TAILBOOM,RIGHTSIDE Avioniccompartment - Check:

CRUand OEU- Condition and security. NOTE

Refer to GNS Operators Manual for systems checks and operation.

TAKEOFF BEFORE

NAV-1/VLFswitch - As desired.

BRGPTR1 switches - As desired. NOTE

Selecting NAV 1, or NAV 2 will connect bearing pointer 1 to either the NAV 1 or NAV 2 receiverrespectively.

BRGPTR2 switches - As desired. NOTE

Selecting VLF or ADF will connect bearing pointer 2 to either the VLF or ADF receiver respectively.

Useor disclosureof data containedonthis pageis subjectto the restrictiononthe title pageof this document.



T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R 4 EW * O Section F VI NLY O Section 3 & 4

BHT-412-FMS-21

GNS-500A/S3 MALFUNCTIONS NOTE

Refer to GNS Operators Manual for system malfunction indications and procedures.

Warning flags

FAIL FLAG

FLAG LOCATION

Navigation

HSI

FAULTCONDITION

CORRECTIVE ACTION

Navigationinformation unreliable.

NAV-1/VLFswitch Pressto change navigation modes.

No Change

Useor disclosure of datacontainedon this pageis subjectto the restrictionon the title pageof this document.

3/4


MODEL

AND 36001 - 36019

SUPPLEMENT FOR CATEGORYA OPERATIONS FOR HELICOPTERSEQUIPPEDWITH PT6T-3BOR PT6T-3BFENGINES VERTICAL TAKEOFF

CERTIFIED 6 JUNE 1986

This supplement shall be attached to Model 412 Flight Manual (BHT-412-FM-2) when operating in Category A conditions and PT6T-3B or PT6T-3BF engines are installed.

The information contained herein supplements information of the basic Flight Manual. For limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual or other applicable supplements.

A DIVISION OF TEXTRON CANADA LTD. ALL RIGHTS RESERVED

6 JUNE REVISION 3 -

1986

24 NOVEMBER 1999


NOTICE PAGE

Additional copies of this publication may be obtained by contacting: CommercialPublication Distribution Center Bell Helicopter Textron P.O. Box 482 Fort Worth, Texas 76101-0482

NP

Rev. 2


LOG OF REVISIONS

Original ....................... 0...................... 6 JUN 86 Revision ...................... 1....................11 NOV 89 Revision ...................... 2....................23 APR 99 Revision ...................... 3....................24 NOV 99

OF PAGES

REVISION NO.

PAGE

PAGE

REVISION NO.

FLIGHT MANUAL

Title............................................................3

NP ..............................................................2

A - B ........................................................ i / ii ............................................................. 1 - 2 ......................................................... 2A /2B .......................................................

3 0 1 1

3.................................................................3

4.................................................................1 4 A / 4B ......................................................1 5 - 10 .......................................................0 11 -- 12 .....................................................1 13 -- 14 .....................................................0 15 / 16 .......................................................0 17 - 20 .....................................................0 21 - 22 .....................................................3 23 / 24 ........................................................3 25...............................................................3

26 - 29 .....................................................1 30 - 31 .....................................................0 32...............................................................

1

33...............................................................0 34 ...............................................................1 33 35 60 .....................................................3 61 /62.3 61 / 62 ........................................................ 3

NOTE


A


BHT-412-FMS-22.2


Original ........................0 Revision ......................1 Revision.......................2 Revision.......................3

...................... 6 JUN 86 ................... 11 NOV 89 ................... 23 APR 99 ................... 24 NOV 99

APPROVED:


Rev. 3

NOV 41999



Introduction

BHT-412-FMS-22.2

INTRODUCTION

A "Category A" takeoff is defined as follows: operation of the helicopter in such a manner that if one engine fails at any time after the start of a takeoff, the helicopter can: 1. Return to, and safely stop on, the takeoff area; or

2. Continue the takeoff, climb out, and attain single engine forward flight.

A "Category A" landing is defined as follows: Operation of the helicopter in such a manner that if one engine fails at any point in the approach, the helicopter can 1. Land, and stop safely on the intended landing area; or

2.

Climb out from the point of failure and attain single engine forward flight.

This supplement is divided into

limitations,

procedures, and

i/ii



Section 1

BHT-412-FMS-22.2

ATTENTION

Mandatory compliance with the operating limitations Manual is required by law.

WEIGHT LIMITATIONS

Maximum Gross Weight 10,500 Pounds (4762.7 Kilograms).

TAKEOFF AND LANDING WEIGHT vs ALTITUDE LIMITATIONS

Maximum Takeoff and Landing Weight - Varies with temperature and altitude - See Gross Weight Limits for Takeoff and Landing Chart. NOTE: The minimum heliport width and length are 72 feet (22 meters) and 150 feet (46 meters) respectively for Category "A" vertical operations from ground level or elevated heliports.

ALTITUDELIMIT FOR TAKE-OFF AND LANDING

The altitude limit for takeoff and landing is 2500 feet pressure altitude.

ALTIMETER(VERTICAL T. O. ALTIMETER)

For vertical takeoff procedures the Vertical takeoff altimeter must be set, 100% N2, RPM-flatpitch, doors and windows closed, heater and vent off. NOTE Doors and windows remain closed, heater off, vent off until CDP is reached. This is required to prevent possible errors in the Vertical takeoff altimeter. Rev. 1

1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412ROTORCRAFT FLIGHT MANUAL

Section 1


BHT-412-FMS-22.2

VERTICAL

OPERATIONS

10.600

9.800

9 400

\\ \\

9,000

8.600

\\

8,200

7,800

7.400

30

20

10 5 0

HEADWIND

-

KT

0

5

10

15

20

25

OAT-

°C

30

35

40

45

NOTE SEE HEADWIND COMPONENT CHART, SECTION 4.

GROSS WEIGHT LIMITS FOR TAKEOFFAND LANDING (ENGLISH) 2

Rev. 1



Section 1

BHT-412-FMS-22.2

VERTICAL OPERATIONS

4,800

CHART, SECTION 4.

GROSSWEIGHT LIMITS FOR TAKEOFF AND LANDING (METRIC)

412099-13

Rev. 1

2A/2B



Section 1

BHT-412-FMS-22.2

AIRSPEEDLIMITATIONS

No change to basic envelope. For takeoff and landing airspeed refer to Performance Section.

AMBIENT AIRTEMPERATURE - OPERATING LIMITATIONS See Performance Section in this Supplement.

CROSSWIND LIMITATIONS

The cross wind limit is that combination of wind velocity and direction where the cross wind component exceeds 15 knots. Refer to "Headwind Component Chart" in Section 4.

TYPEOF OPERATION

Category A operation is approved for day/night VMC, non-icing conditions. Night takeoff and landing may be accomplished with adequate lighting.

FLIGHTCREW

The minimum crew for vertical type takeoff and vertical type landing operations consists of two pilots.

CONFIGURATION

Skid landing gear only. All doors on. Pratt and Whitney PT6T-3B or PT6T-3BF engines shall be installed.

ONLY OPERATIONS LIMITS- FORVERTICAL CENTEROFGRAVITY

Center of gravity limits are from Station 130.0 (3302.0) to Station 142.0 (3606.8). The center of gravity operational range is variable, depending upon gross weight and shall be computed from the weight and balance data.

NOTE Station 0 (datum) is located 20 inches (508.0 millimeters) aft of the most forward point of the cabin nose Maximum asymmetric center of gravity limits are 3.5 inches (88.9 millimeters) to the left and right from the fuselage center line.

GENERATOR AMMETER

Maximum = 150 Amps per Ammeter

Rev. 3

3


Section 1


BHT-412-FMS-22.2

EQUIPMENT REQUIRED

This supplement requires the installation of an approved copilots instrument dual control kit, the 212-706-029 altimeter, and one operative SCAS.

11000

LONGITUDINAL C.G. STATION- INCHES

CENTER OFGRAVITY LIMITS (ENGLISH) 4

Rev. 1

kit, an approved



Section 1

BHT-412-FMS-22.2

4,800

4,600

4,400

LONGITUDINAL C.G. STATION - MM.

CENTEROF GRAVITY LIMITS (METRIC)

412099-14

CENTEROF GRAVITY LIMITS (METRIC) Rev. 1

4A/4B


412ROTORCRAFT FLIGHT MANUAL

Section 2

BHT-412-FMS-22.2

Power Assurance Check - Refer to Section 4 of this Supplement.

VERTICAL TYPE TAKEOFF

NOTE

See Vertical Takeoff Profile, Vertical Takeoff Figure 1, Vertical Takeoff Figure 2 and Vertical Takeoff Figure 3. NOTE

Takeoff will be initiated with the helicopter positioned such that the takeoff index marks are directly opposite the crew doors and the helicopter centered on the heliport. This will assure that the tail rotor is within the confines of the heliport.

Triple Tachometer - 100%

Collective - Flat Pitch (full down). Doors and Windows - Closed

Heater and Vent Blower - Off During Takeoff

Vertical Takeoff Altimeter - Set to zero.

Flight Altimeter - Set to correct station pressure or elevation.

In Hover at Two to Four Feet - Note Transmission Torque.

Collective - Apply smoothly to obtain a steady rate of climb along the takeoff flight path using a transmission torque not to exceed an additional 15% torque in excess of the value noted in hover. 5


BHT-412-FMS-22.2

CDP

160 FT. REF. FIG.



FAA APPROVED

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL

Section 2

BHT-412-FMS-22.2

VERTICAL TAKEOFF FIGURE 1

7



2


8


T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHT MANUAL

Section 2

BHT-412-FMS-22.2


9


The pilot will control the flight path by visual reference to the far right corner of the heliport and the takeoff index marks. As altitude above the heliport is increased to approximately 40 to 60 feet, it will be necessary to transition into rearward flight at a very slow speed in order to maintain visual reference with the far right corner of the heliport (see Vertical Takeoff Figure 1). The takeoff will be continued by visual reference to the far right corner of the heliport (see Vertical Takeoff Figure 2) until the critical decision point (CDP) is reached (see Vertical Takeoff Figure 3). NOTE

Visual reference with the heliport is defined as that position where the far right corner of the heliport is aligned approximately halfway between the edge of the instrument panel and the lower corner of the windshield. The amount of heliport area visible to the pilot will vary with height above the ground. At all points in the takeoff maintain visual contact with forward right hand corner of the heliport boundary.

Copilot will call out the following altitudes from the vertical takeoff altimeter during takeoff: a.

At 60 feet indicated: "60 feet"

b.


c.


d.


e.

At 140 feet indicated:"140 feet"

f.

At 160 feet indicated: "CDP, ROTATE"

At the indicated Critical Decision Point (CDP), translate into forward flight to obtain takeoff safety speed (VTOSS) of 30 knots plus wind (65 knots maximum). Apply power of not less than73% torque and climb to 200 feet above takeoff point at VTOSS. Accelerate to best rate of climb airspeed (65 knots) and climb en route. Copilot should monitor power and systems parameters.



Section 2

BHT-412-FMS-22.2

VERTICAL TYPE LANDING

NOTE See Vertical Landing Profile, Vertical Landing Figure I and Vertical Landing Figure 2.

A vertical type landing is a landing that is initiated from the Landing Decision Point (LDP) which is 30 knots indicated airspeed plus reported wind velocity and at an altitude of 200 feet above the level of the heliport surface. Flight Controls - Adjust frictions to desired levels. Governor Switches - Automatic

Twist Grip N1 Controls - Full Open - Throttle Frictions adjusted to desired level.

Engine RPM - 100% N2.

Force Trim - As desired.

Flight Altimeter - Set to nearest reporting station

From an airspeed of 30 knots indicated plus wind speed and a height of 200 feet, the pilot will initiate the approach when the LDP is reached. The LDP is reached when the pilot obtains the correct sight picture of the heliport (see Vertical Landing Figure 1). Pilot calls "LDP" and indicates the approach.

The approach is initiated by raising the nose of the aircraft to obtain the correct approach sight picture (see Vertical Landing Figure 2) and simultaneously lowering the collective to establish the approach angle.

During the descent, visual contact is maintained with the forward right hand corner of the heliport (see Vertical Landing Figure I and Vertical Landing Figure 2). NOTE

The approach angle is such that the tail rotor will clear a 25-foot (7.6-meter) obstacle on the approach end of the heliport.

Rev. 1

11

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O


Section 2

BHT-412-FMS-22.2

VERTICAL

FIGURE

LANDING 1

13

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Section 2


BHT-412-FMS-22.2

VERTICAL LANDING FIGURE 2

14



Section 2

BHT-412-FMS-22.2

As the helicopter crosses the approach end of the heliport with the required tail rotor obstacle clearance, a slight flare is initiated so that the helicopter is brought to a landing with the pilot's door even with or slightly forward of the takeoff mark on the heliport. NOTE

The copilot will call out airspeed and altitude information prior to LDP (i.e., "Airspeed High" or "Airspeed Low," and "Altitude High" or "Altitude Low") and rotor rpm, torque, and Inter Turbine Temperature (ITT) during the approach from the LDP. In the event of an engine failure after LDP, the pilot will adjust the power as soon as possible to obtain the maximum single engine power available during the descent.

15/16



Section 3

BHT-412-FMS-22.2

PRIOR TO CDP

SEGMENT WORDING

FAULT CONDITION

C BOX OIL PRESS

Combining gearbox oil pressure below normal.

C BOX OIL TEMP

Combining gearbox Land oil temperature above limit.

XMSNOIL TEMP

Transmission oil temperature is above limit.

Land

XMSNOIL PRESS

Transmission oil pressure is low.

Land

BAGGAGEFIRE


Land

FIRE I PULL

Fire indication in No. I Engine compartment.

Land, then pull No. 1handle, select MAIN bottle, then RESERVE if necessary, close No. 1 twist grip.

FIRE 2 PULL

Fire indication in No. 2 engine compartment.

Land, then pull No. 2 handle, select MAIN bottle, then RESERVE if necessary, close No. 2 twist grip.

CORRECTIVEACTION Land




BHT41 2-FMS-22.2

PRIOR TO CDP(Cont)

SEGMENT WORDING

FAULT CONDITION

ENG 1 OUT

No. I Engine N1 See Engine Out Procedure. RPM abnormally low.

ENG 2 OUT

No. 2 Engine N1 See Engine Out Procedure. RPM abnormally low.

CORRECTIVE ACTION

AFTER CDP

SEGMENT WORDING

FAULT CONDITION

C BOX OIL PRESS


C BOX OIL TEMP

Accelerate to VTOSS and Combining gearbox oil temperature above land as soon as practicable. limit.

XMSNOIL TEMP

Transmission oil Temperature is above limit.

Accelerate to VTOSS and land as soon as practicable.

XMSN OIL PRESS



BAGGAGEFIRE



FIRE 1 PULL

Fire indication in No. 1Engine compartment.

Accelerate to VTOSS and then pull No. 1handle, select MAINbottle, then RESERVE if necessary, close No. I twist grip.

CORRECTIVE ACTION




Section 3

BHT-412-FMS-22.2

AFTER CDP(Cont)

SEGMENT WORDING

FAULT CONDITION

FIRE 2 PULL

Fire indication in No. 2 Engine compartment.

ENG I OUT

No. I Engine N1 See Engine Out Procedure. RPM abnormally low.

ENG 2 OUT

No. 2 Engine N1 See Engine Out Procedure. RPM abnormally low.

CORRECTIVE ACTION

Accelerate to VTOSS and then pull No. 2 handle, select MAIN bottle, then RESERVE if necessary, close No. 2 twist grip.

PRIOR TO LDP

SEGMENT WORDING

FAULT CONDITION

C BOX OIL PRESS


Reduce power. Land as soon as practicable.

C BOX OIL TEMP


Reduce power. Observe temperature within limits, if not land as soon as possible.

XMSNOIL TEMP



XMSN OIL PRESS



BAGGAGE FIRE


Reduce power to minimum required. Land as soon as possible, and inspect tail boom area for damage.

CORRECTIVEACTION

19




BHT-412-FMS-22.2

PRIOR TO LDP(Cont)

SEGMENT WORDING

FAULT CONDITION

FIRE 1 PULL

Fire indication in No. 1Engine compartment

Pull No. I handle, select MAIN bottle, then RESERVE if necessary, close No. I twist grip.

FIRE 2 PULL

Fire indication in No. 2 Engine compartment

Pull No. 2 handle, select MAINbottle, then RESERVE if necessary, close No. 2 twist grip.

ENG 1 OUT

See Engine Out Procedure. No. I Engine N1 RPMabnormally low.

ENG 2 OUT

No. 2 Engine N1 See Engine Out Procedure. RPMabnormally low.

CORRECTIVEACTION

AFTERLDP

20

SEGMENT WORDING

FAULT CONDITION

C BOX OIL PRESS


C BOX OIL TEMP

Land. Combining gearbox oil temperature above limit.

XMSN OIL TEMP


Land

XMSN OIL PRESS


Land.

BAGGAGEFIRE


Land.

CORRECTIVEACTION Land.



Section 3

BHT-412-FMS-22.2

AFTERLDP (Cont)

SEGMENT WORDING

FAULT CONDITION

FIRE 1 PULL



FIRE 2 PULL



ENG 1 OUT

No. 1 Engine N1 RPM abnormally low.

See Engine Out Procedure.

ENG 2 OUT



CORRECTIVE ACTION

ENGINE OUTPROCEDURE NOTE

After an engine failure, the power on the remaining engine shall be increased to the maximum permissible power limits (2.5 minute power rating for helicopters equipped with PT6T-3B engines or 30 minute power rating for helicopters equipped with PT6T-3BF engines). The rotor speed shall be maintained within limits.

DuringTakeoffPriorto CriticalDecisionPoint(CDP)

An engine failure prior to reaching CDP (160 feet above the heliport) will necessitate a landing back to the heliport. The landing is accomplished by descending back toward the takeoff surface, while maintaining 97% rotor speed. While maintaining a level attitude, increase collective pitch as necessary to cushion the landing. Perform normal shutdown procedure.

Rev. 3

21


Section 3


BHT-412-FMS-22.2

DuringTakeoffAfterCriticalDecisionPoint(CDP)

If an engine fails during or following CDP, the helicopter shall be accelerated to takeoff safety speed (VTOSS) of 30 knots plus wind. When accelerating to VTOSS, rotor speed shall be maintained within limits. Accomplish climbout at an airspeed of 30 knots plus headwind. Use the 2.5 minute power rating (for helicopters equipped with PT6T-3B engines) or the 30 minute power rating (for helicopters equipped with PT6T-3BF engines) at 200 feet above the heliport. Then accelerate to the best rate of climb speed (65 knots). The copilot should monitor power. Shut down affected engine. Twist Grip - Rotate to full closed. Fuel - OFF

Boost Pump - Off

Crossfeed - Override Close Interconnect

Valve - Open

DuringLandingPrior to LandingDecision Point (LDP)

An emergency condition during landing prior to the LDP (200 feet above the heliport), the helicopter should be accelerated to best rate-of-climb speed for climbout, depending on the terrain and obstacles. Shut down affected engine. Twist Grip - Rotate to full closed Fuel - Off

Boost Pump - Off

Crossfeed - Override Close Interconnect

Valve - Open

OR, proceed to the LDP and use the procedure below.

22

Rev. 3



Section 3

BHT-412-FMS-22.2

Point(LDP) attheLanding Decision During Landing

Maintain rotor speed within limits and accomplish a climbout or accelerate rate-of-climb speed of 65 knots, depending on the terrain and obstacles.

to the best

OR, proceed to a landing while using the procedure below.

Decision Point(LDP) AftertheLanding DuringLanding

If an emergency occurs after LDP, the helicopter is committed to land. The landing is accomplished using the 2.5 minute power rating (for helicopters equipped with PT6T-3B engines) or the 30 minute power rating (for helicopters equipped with PT6T-3BF engines). The rotor speed shall be maintained within limits. Visual contact with the sight picture of the two sides of the heliport shall be maintained. After landing, perform normal shutdown procedures.

MALFUNCTION PROCEDURES

Section 3

No change.

Rev. 3

23/24



Section 4

BHT-412-FMS-22.2

The performance data presented in this section are based on the engine manufacturer's specification power for the PT6T-3B or PT6T-3BF engines with installation losses.

minimum

I

An engine power assurance check chart is presented which is also based on a minimum specification engine with installation losses and proper rigging of engine controls.

If engine performance does not meet that shown in the "ENGINE POWER ASSURANCE CHART" steps should be taken to ascertain the causes of engine power loss.

The minimum heliport size, using the vertical takeoff and landing procedure, is 72 feet x 150 feet (22 meters x 46 meters). A chart is presented showing the MINIMUM HELIPORT SIZE AND AN APPROVED HELIPORT MARKING. The heliport marking shown was used during the type tests. The gross weight limits for takeoff and landing varies with the component of head wind directly opposed to the flight path. The Headwind Component Chart presents a method of obtaining the headwind component for use on the Gross Weight Limits for Takeoff and Landing Chart, Section 1. The maximum cross wind demonstrated was 15 Knots and is shown as a limit on the Headwind Component Chart. Interpolation of all data is allowable but extrapolation is not permitted.

HOVER PERFORMANCE

The Hover Performance chart is presented to show the percent torque required to hover in-ground-effect (IGE) at a four-foot (1.2 meters) skid height, and the percent torque available as shown in the following example:

Rev. 3

25


Section 4


BHT-412-FMS-22.2

to Hover Required Torque

(OAT)

1.

Determine Ambient Air Temperature

2.

Determine Pressure Altitude (Hp)

3.

Determine Gross Weight (G.W.)

4.

Enter chart at OAT, proceed vertically upward to Hp proceed horizontally to the right to G.W., then proceed vertically downward to the TORQUEMETER scale and read percent torque required to hover (Transmission Torque).

29°C

1000 ft.

8500 lb. (3855.6 kg.)

59.5%

(TwinEngine) Available Torque 1.

Enter chart at OAT, proceed vertically upward to Hp proceed horizontally to the right to the TAKEOFF POWER AVAILABLE curves, then proceed vertically downward to the TORQUEMETER scale and read percent torque available.

100%

PERFORMANCE TAKEOFF TakeoffSafetySpeed(Vtoss)

The takeoff safety speed varies with wind (Vtoss = 30 kt. IAS + wind).

Calibration Altimeter

Takeoff performance is based on the altimeter calibration shown in the Takeoff Altimeter Calibration Chart.

Takeoff Distance

I

Vertical Takeoff Distance.

Using the vertical takeoff procedure, the takeoff distance measured from the aft end of the heliport when in takeoff position, is the maximum

26

Rev. 1

150 ft. (46 m.)



Section 4

BHT-412-FMS-22.2

distance needed to land after a rejected takeoff (one engine failure prior to or at critical decision point for the vertical takeoff procedure). This takeoff distance applies to all conditions of GROSS WEIGHT, PRESSURE ALTITUDE, AND AMBIENT AIR TEMPERATURE When operating within allowable limits.

TakeoffFlightPath,ObstacleClearance

The takeoff flight path begins at the end of the takeoff distance, at 35 feet above the takeoff surface and Vtoss. Two charts are involved in the determination of the takeoff flight path and are titled as below:

Takeoff Flight Path, Climb Index

Takeoff Flight Path, Obstacle Clearance.

These charts are used as in the following example.

Example

1. Determine Ambient Air Temperature

2.

15°C

Determine Pressure Altitude

Sea Level (0 ft)

3. Determine VTOSS (30 kt. IAS + Headwind Component 40 kts. of 10 Knots)

4.

Determine Actual Gross Weight

8300 lb. (3765 kg.)

5. Determine height above and distance from the Takeoff Surface of known obstacles along the Flight Path.

Rev. 1

27


Section 4


BHT-412-FMS-22.2

6.

Check Maximum Allowable Gross Weight

10,500 lb. (4762.7 kg.)

If actual gross weight is less than allowable, proceed.

7.

Enter Takeoff Flight Path, Climb Index Chart at Ambient Air Temperature.

8.

Move Vertically up to Pressure Altitude.

9.

Move Horizontally Right to Actual Gross Weight for Takeoff.

10. Move Vertically Down to VTOSS Correction Curves. 11. Move Diagonally, Parallel to VTOSS Correction Curves to VTOSS.

12. Move Vertically Down to Climb Index Scale and Read Climb Index.

46.0

13. On the TAKEOFF FLIGHT PATH OBSTACLE CLEARANCE CHART locate CLIMB INDEX point, which has just been determined, at a height above takeoff surface of 200 feet. A line from this point through 35 foot height at 0 Horizontal Distance represents the minimum height flight path from the end of the takeoff distance and should be compared with the height of known obstacles along the flight path for obstacle clearance.

CLIMB PERFORMANCE

SingleEngineat MinimumVtoss(30 kts IAS)

Single engine rate of climb at minimum VTOSS (30 IAS) is shown for 7000, 8000, 9000, 10,000, and 10,500 pounds gross weight. In addition, chart is shown for 3200, 3600, 4000, 4400, and 4762.7 kilograms gross weight. These curves are for general information only, since takeoff flight path chart is presented for flight path determination.

28

Rev. 1

T N I R P N O R I O T F C U T D O R E P E R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

Single Engine at Best Rate of Climb Speed165kts Vcal)

Single engine rate of climb at best rate of climb speed(65 kts Vcal) is unchanged from the basicRotorcraft Flight Manual Data (Cat. B).

LANDING

PERFORMANCE

Landing Distance.Vertical Procedure: Actual Landing Distance ScheduledLanding Distance

110 ft. (33.5m.) 133 R. (42.0 m.) I

Using the vertical procedure,the landing distance is the actual distance neededfor the tail rotor to clear a 25 foot (8 meter) height and cometo a stop on the landing surface with only one engine operating.

The actual landing distance is 110 feet (33.5 meters). For scheduled landings the landing distance is 133 feet (42.0 meters). These landing I distances apply to all conditions of GROSS WEIGHT, PRESSURE ALTITUDE, AND AMBIENT TEMPERATURE when operating within allowable limits.

Rev. 1

29


T N I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O Section 4


BHT-412-FMS-22.2

MODEL 412 POWER ASSURANCE CHECK (HOVER) PT6T-3B ENGINE

HEATER/ECU -

OFF.

THROTTLES: TEST ENGINE - FULL OPEN,FRICTIONED. OTHER ENGINE - FLIGHT IDLE. Nil RPM -

97%.

COLLECTIVE PITCH - INCREASE UNTIL LIGHT ON ° SKIDS OR HOVERING.DO NOT EXCEED810 ITT OR 100.8% NI RPM.

500

550

MAXIMUM 600

STABILIZE POWER ONE MINUTE, THEN RECORD PRESSUREALTITUDE, OAT, ENGINE TORQUE, ITT, AND NI RPM.

IF INDICATED ITT OR NI RPM EXCEEDS MAX ALLOWABLE, REPEATCHECK, STABILIZING POWER FOUR MINUTES. REPEATCHECKUSING OTHERENGINE.

ENTERCHARTAT INDICATEDENGINETORQUE,MOVE UP TO INTERSECTPRESSUREALTITUDE, PROCEED TO THE RIGHT TO INTERSECT OUTSIDE AIR TEMPERATURE,THEN MOVE UP TO READ VALUES FORMAXIMUM ALLOWABLE ITT AND NI RPM.

ALLOWABLE ITT - °C 650 700

IF EITHER ENGINE EXCEEDSALLOWABLE ITTOR NI RPMAFTERSTABILIZNGFOURMINUTES, PUBLISHED PERFORMANCEMAY NOT BE ACHIEVABLE. CAUSE SHOULD BE DETERMINEDAS SOON AS PRACTICAL.

MAXIMUM

800

750

80 60 70 40 50 ENGINE TORQUE - PERCENT (INDICATED)

30

85

ALLOWABLE NI RPM 90

95

PERCENT 100

105

412099-1D



Section 4

BHT-412-FMS-22.2

260

240

220

0

20

CRITICAL

DECISION POINT

40

80

INDICATED

60

TAKEOFF

100

120

ALTIMETER

140 160 -

180 200

FT.

TAKEOFF ALTIMETER CALIBRATION

31


T N I R P N O R I O T F C U T O OD N R i E P E E R R C S N G O N I R W O E I F V NLY O Section 4

BHT-412-FMS-22.2

I

1 FT.(TYP.)

(305 MM.

-

TYP.)

4 FT.X 4 FT. X 4 FT. (TYP.)

(1220 MM. X 1220 MM. X 1220 MM. TYP.)

4 FT.(TYP.) (1220 MM. TYP)

8 FT.(TYP.)

(2440MM. TYP.)

72 FT. (22 M.)




Section 4

BHT-412-FMS-22.2

HEADWIND COMPONENT- KNOTS 15 C 20

25

30

35

40

CROSSWIND COMPONENT LIMIT WIND DIRECTION FROM TAKE OFF

°

2. 3 4. 5. 6

REPORTEDWIND DIRECTION..... 200 ° WIND DIRECTION. DEGREESFROM TAKE OFF HEADING ............. 30 REPORTEDWIND SPEED. ..... . . . 20 KNOTS ENTER CHART AT REPORTEDWIND SPEED.POINT A PROCEEDUPWARD,FOLLOWING THE SHAPEOF THE CURVED LINES, TO WIND DIRECTION. DEGREESFROM TAKE OFF HEADING. POINT B. 7. PROCEEDVERTICALLY UPWARD TO THE HEADWIND COMPONENTSCALE AND READ HEADWIND COMPONENT .......... 17 5 KNOTS

HEADWINDCOMPONENTCHART

33


Section 4


BHT-412-FMS-22.2

HOVER PERFORMANCE

(ENGLISH) IGE (4 FT. SKID HT.) POWERREQUIREDTO HOVER AND TAKE-OFF POWERAVAILABLE ENGINE RPM 100% GENERATOR 150 AMPS

HEATER OFF NO WIND



Section 4

BHT-412-FMS-22.2

HOVER PERFORMANCE

IGE (1.2 M SKIDHT.) POWERREQUIREDTO HOVERAND TAKE-OFFPOWERAVAILABLE

ENGINE RPM100% GENERATOR150AMPS

HEATER OFF NO WIND


Section 4


BHT-412-FMS-22.2

ENGINE PT6T-3B TAKEOFF FLIGHT PATH (ENGLISH) CLIMB INDEX

2.5MINUTE POWER ENGINE RPM 97% GENERATOR 150 AMPS

HEATER OFF INOPERATIVE ENGINE SECURED AIRSPEED = VTOSS (30 KT. IAS + HEADWIND)

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT FLIGHTMANUAL


Section 4

BHT-412-FMS-22.2

PT6T-3B ENGINE TAKEOFF FLIGHT PATH (METRIC) INDEX CLIMB

2.5 MIN. POWER ENGINERPM97% GENERATOR150 AMPS

HEATEROFF INOPERATIVEENGINESECURED AIRSPEED= VTOSS(30KT.IAS + HEADWIND)


Section 4


BHT-412-FMS-22.2

PT6T-3B ENGINE PATH FLIGHT TAKEOFF

(ENGLISH) OBSTACLE CLEARANCE

2.5MINUTE POWER ENGINE RPM 97% GENERATOR 150AMPS

200

HEATER OFF INOPERATIVE ENGINE SECURED AIRSPEED VTOSS = 30 KT. IAS + HEADWIND

30

CLIMB INDEX 25 20



Section 4

BHT-412-FMS-22.2

PT6T-3BENGINE TAKEOFFFLIGHTPATH (METRIC) OBSTACLE CLEARANCE

2.5 MIN. POWER ENGINERPM 97% GENERATOR150 AMPS

HEATEROFF INOPERATIVEENGINESECURED AIRSPEEDVTOSS = 30 KT.IAS = HEADWIND


Section 4


BHT-412-FMS-22.2

PT6T-3BENGINE SINGLEENGINERATEOFCLIMB (ENGLISH) WEIGHT 7,000LB.GROSS

2.5MIN. POWER ENGINE RPM 97% 150 AMPS GENERATOR

3,000

2,000

AIRSPEED VTOSS = 30 KT. IAS HEATER OFF INOPERATIVE ENGINE SECURED


2.5 MIN. POWER ENGINE RPM 97% GENERATOR 150 AMPS

3.000

2.000

1.000



Section 4


BHT-412-FMS-22.2

PT6T-3BENGINE SINGLEENGINERATEOF CLIMB (ENGLISH) WEIGHT 8,000LB.GROSS

2.5MIN. POWER ENGINE RPM 97% GENERATOR 150 AMPS

3,000

2,000


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 0

-1,000

-2,000



4


BHT-412-FMS-22.2

PT6T-3B ENGINE SINGLE ENGINE RATEOFCLIMB (ENGLISH) 9,000LB.GROSS WEIGHT


3,000

2,000




Section 4

BHT-412-FMS-22.2

ENGINE PT6T-3B RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 4000KG.GROSS


3,000

-2.000


I


Section 4


BHT-412-FMS-22.2

PT6T-3BENGINE SINGLEENGINERATEOFCLIMB (ENGLISH) WEIGHT LB.GROSS 10,000


3,000 -

2,000

1,000 -

-1,000 -

-2,000

0

400

AIRSPEED VTOSS = 30 KIAS HEATER OFF INOPERATIVE ENGINE SECURED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 2.5 MIN. POWER ENGINE RPM 97% GENERATOR 150 AMPS



Section 4


BHT-412-FMS-22.2

ENGINE PT6T-3B RATEOFCLIMB ENGINE SINGLE

10,500LB.(4762.7KG.)GROSSWEIGHT

2.5MIN. POWER ENGINE RPM 97% 150 AMPS GENERATOR




Section 4

BHT-412-FMS-22.2

ENGINE PT6T-3BF PATH FLIGHT TAKEOFF

30 MINUTE POWER ENGINE RPM 97% GENERATOR 150 AMPS

(ENGLISH) CLIMB INDEX



BHT-412-FMS-22.2




ENGINE PT6T-3BF PATH FLIGHT TAKEOFF


PS

10080 60 50

40


30

CLIMB INDEX 20 25

18

16

14

12

HORIZONTALDISTANCE- FT.(FROMFORWARDEDGEOFTAKEOFFSURFACE)

Rev. 3

51


Section 4

BHT-412-FMS-22.2

ENGINE PT6T-3BF TAKEOFFFLIGHTPATH (METRIC) OBSTACLE CLEARANCE

30 MINUTE POWER ENGINE RPM 97%

I

U

HEATEROFF INOPERATIVEENGINESECURED




Section

4

BHT-412-FMS-22.2

ENGINE PT6T-3BF RATEOFCLIMB ENGINE SINGLE (ENGLISH) WEIGHT LB.GROSS 7,000



3,000

Rev. 3

53



4


BHT-412-FMS-22.2

PT6T-3BFENGINE SINGLEENGINERATEOF CLIMB (METRIC) WEIGHT 3,200KG.GROSS


2.000




Section 4

BHT-412-FMS-22.2

PT6T-3BFENGINE SINGLEENGINERATEOFCLIMB (ENGLISH) WEIGHT 8,000LB.GROSS

30 MINUTE POWPER ENGINE RPM 97% GENERATOR 150 AMPS


3,000

2,000

RATE OF CLIMB - FT/MIN.

Rev. 3

55


Section 4

BHT-412-FMS-22.2

ENGINE PT6T-3BF SINGLEENGINERATEOF CLIMB (METRIC) WEIGHT 3,600KG.GROSS


3,000

2,000

56

Rev. 3





Section 4

BHT-412-FMS-22.2

PT6T-3BF ENGINE RATEOFCLIMB SINGLE ENGINE (ENGLISH) 9,000LB.GROSS WEIGHT

30 MINUTE POWER ENGINE RPM 97% GENERATOR 150AMPS


3,000

-2,000

0

1,600

400

RATE OF CLIMB - FT./MIN.

Rev. 3

57



4

BHT-412-FMS-22.2

ENGINE PT6T-3BF ENGINE RATEOFCLIMB SINGLE (METRIC) WEIGHT 4000KG.GROSS


3,000

2,000

1,000

3





Section 4

BHT-412-FMS-22.2

PT6T-3BFENGINE SINGLEENGINERATEOF CLIMB (ENGLISH) WEIGHT LB.GROSS 10,000



3,000

2,000

-1,000

Rev. 3

59



4


BHT-412-FMS-22.2

ENGINE PT6T-3BF RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 4400KG.GROSS



3,000

2,000

-

0

400

800

1.200

1,600

2,000


412099-21

60

Rev. 3



Section 4

BHT-412-FMS-22.2

PT6T-3BF ENGINE RATEOFCLIMB SINGLE ENGINE

WEIGHT KG.)GROSS 10,500 LB.(4762.7



3,000

2,000

Rev. 3

61/62

T N I R P N O ROTORCRAFT R I O T MANUAL FLIGHT F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-22.3

MODEL

SUPPLEMENT FOR CATEGORYA OPERATIONS

FOR HELICOPTERSEQUIPPEDWITH PT6T-3BEOR PT6T-BGENGINES 36020- 36086 VERTICAL TAKEOFF

This supplement shall be attached to Model 412 Flight Manual (BHT-412-FM-3)when operating in Category A conditions and PT6T-3BEor PT6T-BGengines are installed.

The information contained herein supplements information of the basic Flight Manual.For limitations, Procedures,and Performance Data not contained in this supplement, consult the basic Flight Manualor other applicablesupplements.

NOTICE COPYRIGHT COPYRIGHT 1999 TEXTRONINC. BELL HELICOPTER INC. TEXTRON ANDBELL HELICOPTER CANADALTD. OF TEXTRON A DIVISION ALL RIGHTSRESERVED

REISSUED- 23 APRIL 1999 1999 REVISION1 - 24 NOVEMBER


NOTICE PAGE

Additionalcopiesof this publicationmay be obtainedby contacting: CommercialPublicationDistributionCenter Bell HelicopterTextron P.O.Box 482 Fort Worth, Texas76101-0482

NP


LOG OF REVISIONS

Original ........................ 0 .................... Reissue ........................ 0.................. Reissue ........................0.................. Revision.......................1 ...................

24 JUL 91 25 OCT 95 23 APR 99 24 NOV 99

LOG OF PAGES

REVISION NO.

PAGE

PAGE

REVISION NO.

FLIGHT MANUAL

T itle............................................................1 NP ..............................................................0 A - B .................................................... 1 0 i/ ii ............................................................. 1 - 3 .........................................................0 ................ 1 ................. .................... 4. . 5 - 1 6 .......................................................0 17 18........................................................0 19 - 22 .....................................................0 23 - 24 .....................................................1 25 / 26 ........................................................1 27 ...............................................................1 0 28 - 37 ..................................................... 38 - 62 .....................................................1 63 / 64 ........................................................1

NOTE


A


BHT-412-FMS-22.3


Original ........................ 0 .................... 24 JUL 91 Reissue ........................0.................. 25 OCT 95 Reissue ........................ 0.................. 23 APR 99 Revision ....................... 1 .................. 24 NOV 99

APPROVED:

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT.WORTH, TX 76193-0170 B

Rev. 1

DATE:

NOV2 4 1999


BHT-412-FMS-22.3

INTRODUCTION

A "Category A" takeoff is defined as follows: operation of the helicopter in such a manner that if one engine fails at any time after the start of a takeoff, the helicopter can: 1. Return to, and safely stop on, the takeoff area; or

2. Continue the takeoff, climb out, and attain single engine forward flight.

A "Category A" landing is defined as follows: Operation of the helicopter in such a manner that if one engine fails at any point in the approach, the helicopter can 1. Land, and stop safely on the intended landing area; or

2. Climb out from the point of failure and attain single engine forward flight.

This supplement is divided into limitations, performance for a given set of conditions.

procedures, and

i/ii



ATTENTION

Compliance with the limitations in this section is required by appropriate operating rules.

WEIGHT LIMITATIONS

Maximum Gross Weight 10,500 Pounds (4762.7 Kilograms).

TAKEOFF AND LANDING WEIGHT vs ALTITUDE LIMITATIONS

Maximum Takeoff and Landing Weight - Varies with temperature and altitude - See Gross Weight Limits for Takeoff and Landing Chart. NOTE:

The minimum heliport width and length are 72 feet (22 meters) and 150 feet (46 meters) respectively for Category "A" vertical operations from ground level or elevated heliports.

ALTITUDELIMIT FOR TAKE-OFF AND LANDING

The altitude limit for takeoff and landing is 2500 feet pressure altitude.

ALTIMETER(VERTICAL T. O. ALTIMETER)

For vertical takeoff procedures the Vertical takeoff altimeter must be set, 100%N2, RPM-flatpitch, doors and windows closed, heater and vent off. NOTE

Doors and windows remain closed, heater off, vent off until CDP is reached. This is required to prevent possible errors in the Vertical takeoff altimeter.

1


BHT-412-FMS-22.3

VERTICAL

10,600

10.200

9,800

OPERATIONS



VERTICALOPERATIONS

4,800


BHT-412-FMS-22.3

AIRSPEEDLIMITATIONS

No change to basic envelope. For takeoff and landing airspeed Performance Section.

refer to

LIMITATIONS - OPERATING AIRTEMPERATURE AMBIENT See Performance Section in this Supplement.

LIMITATIONS CROSSWIND

The cross wind limit is that combination of wind velocity and direction where the cross wind component exceeds 15 knots. Refer to "Headwind Component Chart" in Section 4.

TYPEOF OPERATION

Category A operation is approved for day/night VMC, non-icing conditions. Night takeoff and landing may be accomplished with adequate lighting.

FLIGHTCREW

The minimum crew for vertical operations consists of two pilots.

type takeoff and vertical type landing

CONFIGURATION

Skid landing gear only. All doors on. Pratt and Whitney PT6T-3BE or PT6T-3BG engines shall be installed.

LIMITSCENTEROF GRAVITY ONLY OPERATIONS FORVERTICAL

Center of gravity limits are from Station 130.0 (3302.0) to Station 142.0 (3606.8). The center of gravity operational range is variable, depending upon gross weight and shall be computed from the weight and balance data. NOTE Station 0 (datum) is located 20 inches (508.0 millimeters) aft of the most forward point of the cabin nose Maximum asymmetric center of gravity limits are 3.5 inches millimeters) to the left and right from the fuselage center line.

AMMETER GENERATOR

Maximum = 150 Amps per Ammeter

4

Rev. 1

(88.9



REQUIRED EQUIPMENT

This supplement requires the installation of an approved copilots instrument kit, an approved dual control kit, the 212-706-029 altimeter, and one operative AFCS.

11000

I

10000

9000

8000

7000

6000

130

132

134

136

138

140

142

144

LONGITUDINAL C.G. STATION - INCHES

CENTEROF GRAVITY LIMITS (ENGLISH)

5


BHT-412-FMS-22.3

4,800 4,600

4,400

3300

3400

3500

LONGITUDINAL C.G. STATION - MM.

CENTEROF GRAVITY LIMITS (METRIC)

6

3600


BHT-412-FMS-22.3

Power Assurance Check - Refer to Section 4 of this Supplement.

VERTICAL TYPE TAKEOFF

NOTE

See Vertical Takeoff Profile, Vertical Takeoff Figure 1, Vertical Takeoff Figure 2 and Vertical Takeoff Figure 3. NOTE

Takeoff will be initiated with the helicopter positioned such that the takeoff index marks are directly opposite the crew doors and the helicopter centered on the heliport. This will assure that the tail rotor is within the confines of the heliport.

Triple Tachometer - 100%

Collective - Flat Pitch (full down). Doors and Windows - Closed

Heater and Vent Blower - Off During Takeoff

Vertical Takeoff Altimeter - Set to zero.

Flight Altimeter - Set to correct station pressure or elevation. In Hover at Two to Four Feet - Note Mast Torque.

Collective - Apply smoothly to obtain a steady rate of climb along the takeoff flight path using a mast torque not to exceed an additional 15% torque in excess of the value noted in hover.

7


BHT-412-FMS-22.3

VTOSS

CDP

VERTICAL TAKEOFF PROFILE

8

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SUPPLEMENT


T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O VERTICAL TAKEOFF FIGURE 2

10

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-22.3

FAA APPROVED



BHT-412-FMS-22.3

The pilot will control the flight path by visual reference to the far right corner of the heliport and the takeoff index marks. As altitude above the heliport is increased to approximately 40 to 60 feet, it will be necessary to transition into rearward flight at a very slow speed in order to maintain visual reference with the far right corner of the heliport (see Vertical Takeoff Figure 1). The takeoff will be continued by visual reference to the far right corner of the heliport (see Vertical Takeoff Figure 2) until the critical decision point (CDP) is reached (see Vertical Takeoff Figure 3). NOTE

Visual reference with the heliport is defined as that position where the far right corner of the heliport is aligned approximately halfway between the edge of the instrument panel and the lower corner of the windshield. The amount of heliport area visible to the pilot will vary with height above the ground. At all points in the takeoff maintain visual contact with forward right hand corner of the heliport boundary.

Copilot will call out the following altitudes from the vertical takeoff altimeter during takeoff: a.


b.


c.


d.


e.

At 140 feet indicated:"140 feet"

f.

At 160 feet indicated: "CDP, ROTATE"

At the indicated Critical Decision Point (CDP), translate into forward flight to obtain takeoff safety speed (VTOSS) of 30 knots plus wind (65 knots maximum). Apply power of not less than 64% torque and climb to 200 feet above takeoff point at VTOSS. Accelerate to best rate of climb airspeed (65 knots) and climb en route. Copilot should monitor power and systems parameters.

12



VERTICAL TYPE LANDING

NOTE

See Vertical Landing Profile, Vertical Landing Figure 1 and Vertical Landing Figure 2.

A vertical type landing is a landing that is initiated from the Landing Decision Point (LDP) which is 30 knots indicated airspeed plus reported wind velocity and at an altitude of 200 feet above the level of the heliport surface. Flight Controls - Adjust frictions to desired levels. Governor Switches - Automatic

Twist Grip N1 Controls - Full Open - Throttle Frictions adjusted to desired level. Engine RPM - 100% N2. Force Trim - As desired

Flight Altimeter - Set to nearest reporting station

From an airspeed of 30 knots indicated plus wind speed and a height of 200 feet, the pilot will initiate the approach when the LDP is reached. The LDP is reached when the pilot obtains the correct sight picture of the heliport (see Vertical Landing Figure 1). Pilot calls "LDP" and indicates the approach. The approach is initiated by raising the nose of the aircraft to obtain the correct approach sight picture (see Vertical Landing Figure 2) and simultaneously lowering the collective to establish the approach angle.

During the descent, visual contact is maintained with the forward right hand corner of the heliport (see Vertical Landing Figure 1 and Vertical Landing Figure 2). NOTE

The approach angle is such that the tail rotor will clear a 25-foot (7.6-meter) obstacle on the approach end of the heliport.

13


BHT-412-FMS-22.3

LOP

200 FT. 30 KTS. WIND * REPORTED REF.FIG. 1 ROTATION COMPLETE REF. FIG. 2

ROTATION

VERTICAL LANDING PROFILE

412900-132

14




16



16



As the helicopter crosses the approach end of the heliport with the required tail rotor obstacle clearance, a slight flare is initiated so that the helicopter is brought to a landing with the pilot's door even with or slightly forward of the takeoff mark on the heliport. NOTE

The copilot will call out airspeed and altitude information prior to LDP (i.e., "Airspeed High" or "Airspeed Low," and "Altitude High" or "Altitude Low") and rotor rpm, torque, and Inter Turbine Temperature (ITT) during the approach from the LDP. In the event of an engine failure after LDP, the pilot will adjust the power as soon as possible to obtain the maximum single engine power available during the descent.

17/18



PRIOR TO CDP

SEGMENT WORDING

FAULT CONDITION

CORRECTIVE ACTION

C BOX OIL

Combining gearbox

Land

PRESS

oil pressure below

normal.

C BOXOIL

Combining gearbox

TEMP

oil temperature above

Land

limit.

XMSNOIL

Transmission oil

TEMP

temperature is above

Land

limit.

XMSN OIL PRESS


Land

BAGGAGE FIRE

Smoke in baggage

Land

compartment.

FIRE I PULL

Fire indication in No. 1 Engine com-

Land, then pull No. 1 handie, select MAIN bottle,

partment.

then RESERVE if neces-

sary, close No. I twist grip.

FIRE 2 PULL

Fire indication in No. 2 engine com-

Land, then pull No. 2 handie, select MAIN bottle,

partment.

then RESERVE if neces-

sary, close No. 2 twist grip.

19

BHT-412-FMS-22.3


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O PRIOR TO CDP(Cont)

SEGMENT WORDING

FAULT CONDITION

ENG 1 OUT



ENG 2 OUT



CORRECTIVE ACTION

AFTER CDP

SEGMENT WORDING

FAULT CONDITION

CORRECTIVE ACTION

C BOX OIL PRESS

Combining gearbox oil pressure below


normal.

C BOX OIL TEMP

Combining gearbox oil temperature above


limit.

XMSN OIL TEMP

Transmission oil Temperature is above limit.


XMSN OIL PRESS



BAGGAGE FIRE



FIRE 1 PULL

Fire indication in No. 1 Engine com-

Accelerate to VTOSS and then pull No. 1 handle,

partment.

select MAINbottle, then RESERVE if necessary, close No. 1 twist grip.

20



AFTER CDP(Cont)

SEGMENT WORDING

FAULT CONDITION

FIRE 2 PULL


ENG 1 OUT

See Engine Out Procedure. No. 1 Engine NI RPM abnormally low.

ENG 2 OUT

See Engine Out Procedure. No. 2 Engine N1 RPM abnormally low.

CORRECTIVEACTION

Accelerate to VTOSS and then pull No. 2 handle, select MAIN bottle, then RESERVE if necessary, close No. 2 twist grip.

PRIOR TO LDP

SEGMENT WORDING

FAULT CONDITION

C BOX OIL PRESS


Reduce power. Land as soon as practicable.

C BOX OIL TEMP



XMSNOIL TEMP



XMSN OIL PRESS



BAGGAGE FIRE


Reduce power to minimum required. Land as soon as possible, and inspect tail boom area for damage.

CORRECTIVEACTION

21


BHT-412-FMS-22.3

PRIOR TO LDP(Cont)

SEGMENT WORDING

FAULT CONDITION

CORRECTIVEACTION

FIRE 1 PULL

Fire indication in No. I Engine compartment

Pull No. I handle, select MAINbottle, then RESERVE if necessary, close No. I twist grip.

FIRE 2 PULL

Fire indication in No. 2 Engine compartment

Pull No. 2 handle, select MAINbottle, then RESERVE if necessary, close No. 2 twist grip.

ENG I OUT


ENG 2 OUT


AFTER LDP

22

SEGMENT WORDING

FAULT CONDITION

C BOX OIL PRESS


C BOX OIL TEMP

Land. Combining gearbox oil temperature above limit.

XMSNOIL TEMP


Land

XMSN OIL PRESS


Land.

BAGGAGEFIRE


Land.

CORRECTIVEACTION Land.



AFTER LDP (Cont)

SEGMENT WORDING

FAULT CONDITION

CORRECTIVE ACTION

FIRE 1 PULL


Land, then pull No. 1 handle, select MAIN bottle, then RESERVE if necessary, close No. 1

twist grip.

FIRE 2 PULL

Fire indication in No. 2

Land, then pull No. 2

Engine compartment.

handle, select MAIN bottle, then RESERVE if necessary, close No. 2

twist grip.

ENG 1 OUT ENG 2 OUT

No. 1 Engine N1 RPM

See Engine Out Proce-

abnormally low.

dure.

No. 2 Engine N1 RPM

See Engine Out Proce-

abnormally low.

dure.

ENGINE OUT PROCEDURE NOTE

After an engine failure, the power on the remaining engine shall be increased to the maximum permissible power limits (2.5 minute power rating for helicopters equipped with PT6T-3BE engines or 30 minute power rating for helicopters equipped with PT6T-3BG engines).

The rotor speed shall be maintained within limits.

DuringTakeoffPriorto CriticalDecisionPoint(CDP)

An engine failure prior to reaching CDP (160 feet above the heliport) will necessitate a landing back to the heliport. The landing is accomplished by descending back toward the takeoff surface, while maintaining 97% rotor speed: While maintaining

a level attitude,

increase collective pitch as

necessary to cushion the landing. Perform normal shutdown procedure.

Rev. 1

23



During Takeoff After Critical Decision Point (CDP)

If an engine fails during or following CDP, the helicopter shall be accelerated to takeoff safety speed (VTOSS) of 30 knots plus wind. When accelerating to

VTOSS,rotor speed shall be maintained within limits. Accomplish climbout at an airspeed of 30 knots plus headwind. Use the 2.5 minute power rating (for helicopters equipped with PT6T-3BE engines) or the 30 minute power rating (for helicopters equipped with PT6T-3BG engines) at 200 feet above the heliport. Then accelerate to the best rate of climb speed (65 knots). The copilot

should monitor power. Shut down affected engine. Twist Grip - Rotate to full closed.

Fuel - OFF

Boost Pump - Off

Crossfeed - Override Close

Interconnect Valve - Open

During LandingPriorto LandingDecision Point (LDP)

An emergency condition during landing prior to the LDP (200 feet above the

heliport), the helicopter should be accelerated to best rate-of-climb speed for climbout, depending on the terrain and obstacles. Shut down affected engine. Twist Grip - Rotate to full closed

Fuel - Off

Boost Pump - Off

Crossfeed - Override Close

Interconnect Valve - Open

OR, proceed to the LDP and use the procedure.

24

Rev.1


BHT-412-FMS-22.3

DuringLandingat theLandingDecisionPoint(LDP)

Maintain rotor speed within limits and accomplish a climbout or accelerate to the best rate-of-climb speed of 65 knots, depending on the terrain and obstacles. OR, proceed to a landing while using the procedure below.

DuringLandingAfterthe LandingDecisionPoint(LDP)

If an emergency occurs after LDP, the helicopter is committed to land. The landing is accomplished using the 2.5 minute power rating (for helicopters equipped with PT6T-3BE engines) or the 30 minute power rating (for helicopters equipped with PT6T-3BG engines). The rotor speed shall be maintained within limits. Visual contact with the sight picture of the two sides of the heliport shall be maintained. After landing, perform normal shutdown procedures.

Rev. 1

25/26



The performance data presented in this section are based on the engine manufacturer's minimum specification power for the PT6T-3BE or PT6T-3BG engines with installation losses. Refer to Basic Flight Manual.

I

The minimum heliport size, using the vertical takeoff and landing procedure, is 72 feet x 150 feet (22 meters x 46 meters). A chart is presented showing the MINIMUM HELIPORT SIZE AND AN APPROVED HELIPORT MARKING. The heliport marking shown was used during the type tests.

The gross weight limits for takeoff and landing varies with the component of head wind directly opposed to the flight path. The Headwind Component Chart presents a method of obtaining the headwind component for use on the Chart, for Takeoff and Landing Gross Weight Limits Section 1. The maximum cross wind demonstrated was 15 Knots and is shown as a limit on the Headwind Component Chart. Interpolation

of all data is allowable but extrapolation

is not permitted.

PERFORMANCE HOVER

The Hover Performance chart is presented to show the percent torque required to hover in-ground-effect (IGE) at a four-foot (1.2 meters) skid height, and the percent torque available as shown in the following example:

Rev.1

27

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAAAPPROVED SUPPLEMENT

BHT-412-FMS-22.3

Torque Requiredto Hover

29°C

1.

Determine Ambient Air Temperature (OAT)

2.

Determine Pressure Altitude (Hp)

3.

Determine Gross Weight (G.W.)

4.

Enter chart at OAT, proceed vertically upward to Hp, proceed horizontally to the right to G.W., then to the vertically downward proceed TORQUEMETER scale and read percent torque required to hover (Mast Torque).

1000 ft.

8500 lb. (3855.6 kg.)

52.3%

Torque Available (Twin Engine) 1.

Enter chart at OAT, proceed vertically upward to Hp, proceed horizontally to the right to the TAKEOFF POWER AVAILABLE curves, then proceed vertically downward to the TORQUEMETER scale and read per cent torque available.

93.0%

TAKEOFF PERFORMANCE Takeoff Safety Speed (Vtoss)

The takeoff safety speed varies with wind (Vtoss = 30 kt. IAS + wind).

Altimeter Calibration

Takeoff performance is based on the altimeter calibration shown in the Takeoff Altimeter Calibration Chart.

Takeoff Distance

Vertical Takeoff Distance

150 ft. (46 m.)

Using the vertical takeoff procedure, the takeoff distance measured from the aft end of the heliport when in takeoff position, is the maximum

28



distance needed to land after a rejected takeoff (one engine failure prior to or at critical decision point for the vertical takeoff procedure). This takeoff distance applies to all conditions of GROSS WEIGHT, PRESSURE ALTITUDE, AND AMBIENT AIR TEMPERATURE When operating within allowable limits.

Takeoff Flight Path, ObstacleClearance

The takeoff flight path begins at the end of the takeoff distance, at 35 feet above the takeoff surface and Vtoss. Two charts are involved in the determination of the takeoff flight path and are titled as below:

Takeoff Flight Path, Climb Index

Takeoff Flight Path, Obstacle Clearance

These charts are used as in the following example.

Example

1.

Determine Ambient Air Temperature

2.

Determine Pressure Altitude

15°C

Sea Level (0 ft)

3. Determine VTOSS (30 kt. IAS + Headwind Component of 10 Knots) 4.

5.

Determine Actual Gross Weight

40 kts.

8300 lb. (3765 kg.)

Determine height above and distance from the Takeoff Surface of known obstacles along the Flight Path.

29



6.

Check Maximum Allowable Gross Weight

10,500lb. (4762.7 kg.)

If actual gross weight is less than allowable, proceed.

7.

Enter Takeoff Flight Path, Climb Index Chart at Ambient Air Temperature.

8.

Move Vertically up to Pressure Altitude.

9.

Move Horizontally Right to Actual Gross Weight for Takeoff.

10. Move Vertically Down to VTOSS Correction Curves. 11. Move Diagonally, Parallel to VTOSS Correction Curves to VTOSS.

12. Move Vertically Down to Climb Index Scale and Read Climb Index.

46.0

13. On the TAKEOFF FLIGHT PATH OBSTACLE CLEARANCE CHART locate CLIMB INDEX point, which has just been determined, at a height above takeoff surface of 200 feet. A line from this point through 35 foot height at 0 Horizontal Distance represents the minimum height flight path from the end of the takeoff distance and should be compared with the height of known obstacles along the flight path for obstacle clearance.

CLIMB PERFORMANCE

SingleEngineat MinimumVtoss(30 kts IAS)

Single engine rate of climb at minimum VTOSS (30 IAS) is shown for 7000, 8000, 9000, 10,000, and 10,500 pounds gross weight. In addition, chart is shown for 3200, 3600, 4000, 4400, and 4762.7 kilograms gross weight. These curves are for general information only, since takeoff flight path chart is presented for flight path determination.



Single Engine at Best Rate of Climb Speed(65 kts Vcal)

Single engine rate of climb at best rate of climb speed (65 kts Vcal) is unchanged from the basic Rotorcraft

Flight Manual Data (Cat. B).

LANDING PERFORMANCE Landing Distance, Vertical Procedure:

Actual Landing Distance .......................... Scheduled Landing Distance .......................

110 ft. (33.5 m.) 138 ft. (42.0 m.)

Using the vertical procedure, the landing distance is the actual distance needed for the tail rotor to clear a 25 foot (8 meter) height and come to a stop on the landing surface with only one engine operating. The actual landing distance is 110 feet (33.5 landings the landing distance is 138 feet (42.0 distances apply to all conditions of GROSS ALTITUDE, AND AMBIENT TEMPERATURE allowable limits.

meters). For scheduled meters). These landing WEIGHT, PRESSURE when operating within

31


BHT-412-FMS-22.3

PAGE INTENTIONALLYLEFT BLANK



260

240 220 200

CRITICAL DECISION POINT

180 160

L

0

140

I

0

20

40

60

80

100

120

140

160

180 200

INDICATED TAKEOFF ALTIMETER - FT.

TAKEOFF ALTIMETER CALIBRATION

33

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 4 FT.X 4 FT. X4 FT. (TYP.) (1220 MM. X 1220 MM. X 1220MM. TYP.)



HEADWIND COMPONENT- KNOTS

CROSSWIND COMPONENT LIMIT WIND DIRECTION FROM TAKE OFF

10 KNOTS

CROSSWIND

COMPONENT

0 KNOTS

REPORTED WINDSPEED- KNOTS

EXAMPLE.

170° ................. . ...... TAKE-OFF HEADING................... 200° REPORTEDWIND DIRECTION .................... 30 .... HEADING.... TAKE-OFF FROM DEGREES WINDDIRECTION, 20 KNOTS . ... . .............. . .... ........ REPORTEDWIND SPEED......... ENTER CHART AT REPORTEDWIND SPEED,POINT A PROCEEDUPWARD, FOLLOWINGTHE SHAPEOF THE CURVED LINES, TO WIND DIRECTION, DEGREESFROM TAKE-OFF HEADING, POINT B 7. PROCEEDVERTICALLY UPWARD TO THE HEADWIND 17.5 KNOTS COMPONENTSCALE AND READ HEADWIND COMPONENT.. .

1. 2 3. 4. 5. 6.

HEADWIND COMPONENT CHART

35




HOVER PERFORMANCE

IGE (1.2 M SKID HT.) POWERREQUIREDTO HOVER AND TAKE-OFFPOWERAVAILABLE

ENGINE RPM100%

HEATER OFF

GENERATOR 150 AMPS

NO WIND


BHT-412-FMS-22.3

PT6T-3BEENGINES TAKEOFFFLIGHTPATH (ENGLISH) CLIMBINDEX

2.5MINUTE POWER ENGINE RPM 97% 150 AMPS GENERATOR

38

Rev. 1

HEATER OFF INOPERATIVE ENGINE SECURED AIRSPEED = VTOSS (30 KT. IAS + HEADWIND)



ENGINES PT6T-3BE FLIGHTPATH TAKEOFF (METRIC) INDEX CLIMB


HEATEROFF INOPERATIVEENGINESECURED AIRSPEED VTOSS(30 KT. IAS + HEADWIND)

o

V

Rev. 1

39


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HT-412-FMS-22.3


ENGINES PT6T-3BE FLIGHTPATH TAKEOFF

(METRIC) CLEARANCE OBSTACLE


HEATEROFF INOPERATIVEENGINESECURED AIRSPEEDVTOSS = 30 KT. IAS = HEADWIND


PT6T-3BE ENGINES SINGLE ENGINE RATEOFCLIMB (ENGLISH) 7,000 LB.GROSS WEIGHT

2.5MIN. POWER ENGINE RPM 97%

AIRSPEED VTOSS = 30 KT. IAS HEATER OFF

3 000

2.000

-2.000

400


42

Rev. 1



ENGINES PT6T-3BE RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 3,200KG.GROSS


3.000

2.000


I


BHT-412-FMS-22.3

PT6T-3BE ENGINES SINGLE ENGINE RATEOFCLIMB (ENGLISH) 8,000LB.GROSS WEIGHT



3,000

2,000

i-


44

Rev. 1



PT6T-3BE ENGINES ENGINE RATEOFCLIMB SINGLE (METRIC) WEIGHT 3,600KG.GROSS


3.000

2.000

AIRSPEEDVTOSS= 30 KT. IAS HEATEROFF INOPERATIVEENGINESECURED


BHT-412-FMS-22.3

PT6T-3BE ENGINES RATEOFCLIMB SINGLE ENGINE (ENGLISH) WEIGHT LB.GROSS 9,000

2.5 MIN. POWER ENGINE RPM 97% 150 AMPS GENERATOR

3.000

2.000

0




ENGINES PT6T-3BE RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 4000KG.GROSS


AIRSPEEDVTOSS= 30 KT. IAS HEATEROFF INOPERATIVEENGINESECURED

2.000

Rev. 1

47


BHT-412-FMS-22.3

ENGINES PT6T-3BE RATEOFCLIMB ENGINE SINGLE (ENGLISH) WEIGHT LB.GROSS 10,000



3.000

2.000


48

Rev. 1



PT6T-3BE ENGINES RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 4400KG.GROSS


VTOSS = 30 KIAS AIRSPEED HEATEROFF INOPERATIVEENGINESECURED

3,000

2,000

1.600

RATEOFCLIMB- FT./MIN.

Rev. 1

49


BHT-412-FMS-22.3

PT6T-3BE ENGINES ENGINE RATEOFCLIMB SINGLE

WEIGHT KG.)GROSS LB.(4762.7 10,500


AIRSPEED VTOSS 30 KT. IAS HEATER OFF INOPERATIVE ENGINE SECURED

3.000

I

1.600


50

Rev. 1

2.000



PT6T-3BG ENGINES TAKEOFF FLIGHTPATH (ENGLISH) CLIMB INDEX


HEATER OFF INOPERATIVE ENGINE SECURED AIRSPEED - VTOSS (30 KT. IAS + HEADWIND)

Rev. 1


BHT-412-FMS-22.3

ENGINES PT6T-3BG FLIGHTPATH TAKEOFF (METRIC) INDEX CLIMB

30 MINUTE POWER ENGINE RPM 97% GENERATOR 150 AMP

HEATEROFF INOPERATIVEENGINESECURED


BHT-412-FMS-22.3

PT6T-3BG ENGINES TAKEOFF FLIGHT PATH




CLIMB INDEX

160

Rev. 1

53


BHT-412-FMS-22.3

PT6T-3BG ENGINES TAKEOFF FLIGHT PATH (METRIC)

30MINUTE POWER ENGINE RPM 97%

CLEARANCE OBSTACLE HEATEROFF

INOPERATIVEENGINESECURED

200

50

100

150

200

250

300

350

400

450

HORIZONTALDISTANCE - METERS(FROMFORWARDEDGEOF TAKEOFFSURFACE)

54

Rev. 1



PT6T-3BG ENGINES SINGLE ENGINE RATEOFCLIMB (ENGLISH) 7,000LB.GROSS WEIGHT




Rev. 1

55


BHT-412-FMS-22.3

PT6T-3BG ENGINES SINGLE ENGINE RATEOFCLIMB (METRIC) 3,200 KG.GROSS WEIGHT



3,000

2,000

1,000

0

r

-1.000

-2,000

0

2.000

400


56

Rev. 1


BHT-412-FMS-22.3

PT6T-3BG ENGINES RATEOFCLIMB ENGINE SINGLE (ENGLISH) 8,000 LB.GROSS WEIGHT



3,000

2.000

w

Rev. 1

57


BHT-412-FMS-22.3

ENGINES PT6T-3BG RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 3,600KG.GROSS


AIRSPEEDVTOSS = 30 KT.IAS HEATEROFF INOPERATIVEENGINESECURED

3.000

2.000

-1.000


58

Rev. 1


BHT-412-FMS-22.3

3.000

2.000

Rev. 1

59


BHT-412-FMS-22.3

ENGINES PT6T-3BG RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 4000KG.GROSS



2,000

1.600

2.000


412099-20

60

Rev. 1



ENGINES PT6T-3BG RATEOFCLIMB ENGINE SINGLE (ENGLISH) WEIGHT LB.GROSS 10,000


AMPS

3,000

800

1,200


Rev. 1

61


PT6T-3BG ENGINES RATEOFCLIMB ENGINE SINGLE (METRIC) WEIGHT 4400KG.GROSS

30 MINUTEOEI POWER

AIRSPEED VTOSS = 30 KIAS

GENERATOR150 AMPS

INOPERATIVE ENGINE SECURED

3,000


62

Rev. 1



PT6T-3BG ENGINES SINGLE ENGINE RATEOFCLIMB

10,500 LB.(4762.7 KG.)GROSS WEIGHT


30 MINUTEOEI POWER ENGINERPM 97% GENERATOR150 AMPS

800

1.200

2.000


Rev. 1

63/64


MODELS


SEAT CUSHIONKIT 412-706-019 CERTIFIED 24 JULY1987

This supplementshall be attached to Model 412 or 412EP Flight Manualwhen the 412-706-019Seat Cushion Kit has beeninstalled. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and PerformanceDatanot containedin this supplement,consult basic Flight Manual.

Bell Helicopter COPYRIGHT NOTICE COPYRIGHT 1995 BELL HELICOPTER INC. ANDBELLHELICOPTER TEXTRON INC. A DIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

A Subsidiary of Textron Inc POST OFFICE BOX 482 · FORT WORTH. TEXAS76101

REISSUE -

8 DECEMBER 1995


NOTICEPAGE

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 76101-0482


LOGOFREVISIONS


0 ...............

24 Jul 87

Reissue...........0

................

8 Dec 95

LOGOF PAGES

PAGE

Title - NP .............................. A - B ..................................

REVISION NO. 0 0

PAGE

i/ii ....................................... 1/2 ......................................

REVISION NO. 0 0

NOTE



FAA APPROVED

LOGOF APPROVEDREVISIONS


...............

24 Jul 87

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE

FEDERAL AVIATIONADMINISTRATION FT. WORTH, TX 76193-0170

B

Reissue ...........

0 ................

8 Dec 95


BHT-412-FMS-24

GENERAL INFORMATION

INTRODUCTION

The Seat Cushion Kit is installed in conjunction with the Utility Passenger Seat Kit and provides increased comfort level for passengers.

i/ii

T N I * Section1 R P N O R I O T F C U T O OD N R E Section P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-24


PLACARDSAND MARKINGS

Actual weight change shall be determined after Seat Cushion Kit is Installed and ballast readjusted if necessary to return empty weight CG within allowable limits.

DOORS MUST BE KEPT CLOSEDDURING FLIGHT IF SEAT CUSHIONSINSTALLED Located on inside of sliding passenger door.

BEFORETAKEOFF

Passengers doors - Closed.

T N I R P N ROTORCRAFT O R I O T FLIGHT MANUAL F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-25.2, 25.3 AND 25.4

MODELS

33108- 33213 36001 - 36019 AND 36020- 36086 AND 36087 AND SUB

SUPPLEMENTFOR AUXILIARY FUEL OPERATIONS (412-706-009) CERTIFIED 10 MARCH 1988

This supplement shall be attached to the 412 Flight Manual (BHT-412-FM-2, -3 or -4) when the 412-706-009 Auxiliary Fuel Kit has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

1994 COPYRIGHT

TEXTRON, HELICOPTER ANDBELL LTD. CANADA OFTEXTRON ADIVISION RESERVED ALLRIGHTS


NOTICE

RIGHTSNOTICE PROPRIETARY

Manufacturer's Data portion of this supplement is proprietary to Bell Helicopter Textron, inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell HelicopterTextron. Inc.

Additionalcopiesof this publicationmaybeobtainedby contacting: DistributionCenter CommercialPublication BellHelicopterTextron PO.Box482 FortWorth,Texas76101


LOG OF REVISIONS

Original ..... Reissue..... Reissue.....

0 0 0

............. ............. .............

10 Mar 88 15 Oct 92 23 Jun 94

LOG OF PAGES

MANUFACTURER'SDATA

FLIGHT MANUAL

PAGE

Title - NP....................... A -- B ..........................

REVISION NO.

1-1 -

0

REVISION NO.

PAGE

1-10

.....................

1-11/1-12 ......................

0

0

i/ii . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

1-1/1-2 ........................ 2-1/2-2 ........................ 3-1/3-2 ......................... 4-1/4-2 .........................

0 0 0 0

NOTE


A


BHT-412-FMS-25.2, 25.3 AND 25.4

LOG OF APPROVEDREVISIONS

Original ..... Reissue ..... Reissue .....

0 0

............ ............. .............

10 Mar 88 15 Oct 92 23 Jun 94

APPROVED:

MANAGER



FAA APPROVED

INTRODUCTION

The Auxiliary Fuel Kit provides additional fuel capacity to extend the range of the helicopter. The kit consists of a left and right auxiliary fuel tank and the hardware and wiring necessary to complete the installation. The left or right auxiliary fuel tank may be removed as operational requirements dictate.

One fuel tank provides an additional 16.3 U.S. gallons (61.7 liters) of fuel. Both fuel tanks combined, provide additional 32.6 U.S. gallons (123.4 liters) of fuel.


BHT-412-FMS-25.2, 25.3 AND 25.4

WEIGHT/CG LIMITATIONS NOTE

The contents of this supplement shall be used in conjunction with the basic Flight Manual for helicopters equipped with the 412-706-009 Auxiliary Fuel Tank Kit installed.

WEIGHT AND BALANCE

Actual weight change shall be determined after kit is installed and ballast readjusted, if necessary, to retain gross weight CG within allowable limits.

FUELAND OIL LIMITATIONS FUEL SYSTEM CAPACITIES

Basic system with left or right auxiliary tank:

Total usable fuel capacity is 346.7 U.S. gallons (1312.3 liters). Basic system with both auxiliary tanks:

Total usable fuel capacity is 363.0 U.S. gallons (13.74.1 liters).

1-1/1-2

T N I Section2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-25.2, 25.3 AND 25.4

IN-FLIGHT OPERATION

CAUTION

WHEN ONE CABIN MOUNTED AUXILIARY FUEL TANK ONLY IS USED, THE TANK INTERCONNECT SWITCH ON THE COCKPIT FUEL PANEL MUST BE PLACED IN THE OPEN POSITION WHEN THE FUEL QUANTITY INDICATIONON THE LOW SIDE WHICH DOES NOT HAVE AN AUXILIARY TANK, REDUCES TO APPROXIMATELY500 LBS. THIS WILL ALLOW THE AUXILIARY FUEL TO BE SHARED BY BOTH ENGINE FEED PRECLUDE THE TANKS AND POSSIBILITY OF FUEL EXHAUSTION TO THE ENGINEBEING SUPPLIEDBY THE SIDE WHICH DOES NOT HAVE THE TANK. AN AUXILIARY AUTOMATIC FEATUREOF THE TANK INTERCONNECT VALVE MAY NOT ONLY ONE FUNCTION WITH AUXILIARY TANK INSTALLED.

2-1/2-2


BHT-412-FMS-25.2, 25.3 AND 25.4


3-1/3-2

T N I Section4 R P N O R I O T F C U 0T O OD N R E P E E R R C S N G O N I W 0 OR E I F Y V L 0 N O FAA APPROVED

BHT-412-FMS-25.2, 25.3 AND 25.4

Noc hange from basic manual.

0

4-1/4-2

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S

DATA

BHT-412-FMS-25.2. 25.3 AND 25.4

MANUFACTURER'S

AUXILIARY FUEL SYSTEM AUXILIARY

FUEL SYSTEM SERVICING

The auxiliary fuel tanks are interconnected withthebasicfuel systemtoallowgravityflow of auxiliary fuel into main fuel cells as fuel is consumed. The auxiliary fuel system is serviced simultaneously with the basic fuel system through the single filler port located on the aft right side of the fuselage. AUXILIARY

FUEL LOADING TABLES

DATA

Fuel loading tables are presented for weight and balance computations in both English and Metric units. These tables shall be used in lieu of the tables for the basic fuel system when either or both auxiliary fuel tanks are installed. Weights and moments listed herein represent total fuel on board to include that contained in basic fuel cells. Refer to table 1-1 and 1-2 for English or 1-1M and 1-2M for Metric when both left and right auxiliary tanks are installed. Tables 1-3 and 1-4 for English or 1-3M and 1-4M for Metric apply to single auxiliary tank installed on left or right.

Total usable fuel capacity with 412-706-009 Auxiliary Fuel Kit (both tanks) installed is 363.0 U.S. gallons (1374.1 liters).

Total usable fuel capacity with one tank (left or right) installed is 346.7 U.S. gallons (1312.3 liters).

1-1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-25.2, 25.3 AND25.4 Table1-1.

MANUFACTURER'SDATA

Fuelloadingwith left andrightauxiliarytanks (two 16.3 gal.) longitudinal(English)

Longitudinal

Jet B orJP-4 (6.5 Lb/U.S. Gallon)

Quantity (U.S. Gal.)

*

*

CG (inches)

Jet A, Al orJP-5 (6.8 Lb/U.S.Gallon)

Moment (In-Lb)

10

65

20 30

130 195

40

139.6 139.8

260

18148 27261

139.9

36374

50 58.3 60

325 379 390

70

139.9 139.9 141.0

455

45468 53022 54990

145.4

66157

139.4

520 585 650 715 780 845 910 975 1040 1105 1122 1170 1235 1300 1338 1365 1430 1495 1560 1625 1690 1755 1792 1820

290

148.4 150.6 152.2 153.5 154.7 155.7 156.5 157.2 157.9 158.6 158.8 156.2 152.7 149.8 148.2 148.7 149.9 151.1 152.2 153.2 154.0 154.8 155.2 154.7

1885

77168 88101 98930 109753 120666 131567 142415 153270 164216 175253 178174 182754 188585 194740 198292 202976 214357 225895 237432 248950 260260 271674 278118 281554

153.2

300 310 320 327.7 330 340

1950 2015 2080 2130 2145 2210

151.9 150.7 149.4 148.7 149.0 149.7

2340 2360

151.2 151.3

360 363.0

2275

150.4

Quantity (U.S. Gal)

9061

80 90 100 110 120 130 140 150 160 170 172.6 180 190 200 205.8 210 220 230 240 250 260 270 275.7 280

350

*

Weight (Pounds)

A

Weight (Pounds)

CG (Inches)

Moment (in-Lb)

10

68

20 30

136 204

40

139.6 139.8

272

18986 28519

139.9

38053

50 58.3 60

340 397 408

139.9 139.9 141,0

476

145.4

47566 55540 57528

70

139.4

9479

69210

80 90 100 110 120 130 140 150 160 170 172.6 180 190 200 205.8 210 220 230 240 250 260 270 275.7 280

288782

544 612 680 748 816 884 952 1020 1088 1156 1174 1224 1292 1360 1399 1428 1496 1564 1632 1700 1768 1836 1875 1904

148.4 150.6 152.2 153.5 154.7 155.7 156.5 157.2 157.9 158.6 158.8 156.2 152.7 149.8 148.2 148.7 149.9 151.1 152.2 153.2 154.0 154.8 155.2 154.7

290

80730 92167 103496 114818 126235 137639 148988 160344 171795 183342 186431 191189 197288 203728 207332 212344 224250 236320 248390 260440 272272 284213 291000 294549

1972

153.2

296205 303661 310752 316731 319605 330837

302110

300 310 320 327.7 330 340

2040 2108 2176 2228 2244 2312

151.9 150.7 149.4 148.7 149.0 149.7

309876 317676 325094 331304 334356 346106

360 363.0

2448 2469

151.2 151.3

370138 373560

342160 353808 357068

*

*

350

NOTE: All data above represents usable fuel (basic and auxiliary) * Quantity at which CG of fuel changes direction.

2380

150.4

based on nominal density

Quantity resulting in maximumforward CGof helicopter(at any weight). Quantity resulting in maximumaft CGof helicopter (weight empty 6520 lb or less). Quantity resulting in maximum aft CG of helicopter (weight empty morethan 6520 lb).

at 15°C

357952

(59°F).


MANUFACTURER'S DATA

Table 1-1M.

Fuel loading with left and right auxiliary tanks (two 61.7 liter) longitudinal (Metric)

Longitudinal

Jet A, Al or JP5 (0.815 kg/liter)


Quantity (liters)

~

40 80 120 160

200

A

220.7 240 280 320 360 400 440 480 520 560 600 640 653.2 680 720 760 778.8 800 840 880 920 960 1000 1040 1043.7 1080 1120 1160 1200 1240.2 1280 1320 1360 1374.1

Weight (kg)

31.2 62.3 93.5 124.6 155.8 171.9 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 498.6 508.8 529.7 560.9 592.0 606.7 623.2 654.4 685.5 716.7 747.8 779.0 810.2 813.0 841.3 872.5 903.6 934.8 966.1 997.1 1028.3 1059.4 1070.4

CG (mm)

3541 3547 3551 3552 3552 3553 3632 3726 3795 3848 3886 3917 3947 3970 3988 4008 4026 4033 3973 3879 3800 3765 3785 3815 3843 3874 3899 3922 3941 3942 3907 3871 3838 3807 3776 3800 3820 3838 3843

Moment (kg-mm)

110479 220978 332019 442579 553402 610761 679184 812641 946094 1078979 1210878 1342748 1475783 1608247 1739566 1873339 2007364 2051990 2104498 2175731 2249600 2284226 2358812 2496536 2634377 2776496 2915672 3055238 3192998 3204846 3286959 3377448 3468017 3558784 3647994 3788980 3928106 4065977 4113547

Quantity (liters)

40 80 120 160 200 220.7 240 280 320 360 400 440 480 520 560 600 640 653.2 680 720 760 * 778.8 800 840 880 920 960 1000 1040 1043.7 1080 1120 1160 1200 • 1240.2 1280 1320 1360 1374.1

Weight (kg)

32.6 65.2 97.8 130.4 163.0 179.9 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 532.3 554.2 586.8 619.4 634.6 652.0 684.6 717.2 749.8 782.4 815.0 847.6 850.5 880.2 912.8 945.4 978.0 1010.7 1043.2 1075.8 1108.4 1119.9

NOTE: All data above represents usable fuel (basic and auxiliary) based on nominal * Quantity at which CG of fuel changes direction. Quantity resulting in maximum forward CG of helicopter (at any weight). Quantity resulting in maximum aft CG of helicopter (weight empty 2957 kg or less). Quantity resulting in maximum aft CG of helicopter (weight empty more than 2957 kg).

CG (mm) 3541 3547 3551 3552 3552 3553 3632 3726 3795 3848 3886 3917 3947 3970 3988 4008 4026 4033 3973 3879 3800 3765 3785 3815 3843 3874 3899 3922 3941 3942 3907 3871 3838 3807 3776 3800 3820 3838 3843

density

Moment (kg-mm)

115437 231264 347288 463181 578976 639185 710419 850273 989736 1129003 1266836 1404636 1544066 1682486 1820123 1959912 2099962 2146766 2201837 2276197 2353720 2389269 2467820 2611749 2756200 2904725 3050578 3196430 3340392 3352671 3438941 3533449 3628445 3723246 3816403 3964160 4109556 4254039 4303776

at 150C (59°F).

1-3


Table 1-2.

MANUFACTURER'S DATA

Fuel loading with left and right auxiliary tanks (two 16.3 gal.) lateral (English)

Lateral

Jet B or JP-4 (6.5 Lb/U.S. Gallon)


*

10 20 30 40 50 54.6 60 70 80.0 90 100 110 120 130 140 150 160 170 172.6 180 190 200 205.8 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 363.0

Weight (Pounds) 65 130 195 260 325 355 390 455 520 585 650 715 780 845 910 975 1040 1105 1122 1170 1235 1300 1337 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2360

CG (Inches) 0 0 0 0 0 0 -0.03 -0.06 -0.05 -0.04 -0.04 -0.03 -0.03 -0.03 -0.03 -0.03 -0.02 -0.02 -0.02 -0.03 -0.44 -0.55 -0.60 -0.59 -0.56 -0.54 -0.52 -0.50 -0.48 -0.46 0.44 -0.43 -0.41 0.40 -0.39 -0.38 -0.36 -0.35 -0.34 -0.34

Jet A, A1 or JP-5 (6.8 Lb/U.S. Gallon)

Moment (In-Lb)

0 0 0 0 0 0 12 -27 -26 -23 -26 -21 -23 -25 -27 -29 -21 -22 -22 -35 -543 -715 -802 -805 -801 -807 -811 813 -811 -807 -801 -811 -800 -806 -811 -815 -796 -796 -796 -802

Quantity (U.S. Gal)

10 20 30 40 50 54.6 60 70 80.0 90 100 110 120 130 140 150 160 170 * 172.6 180 190 200 * 205.8 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 363.0

NOTE: All data above represents usable fuel (basic and auxiliary) * Quantity at which CG of fuel changes direction. Quantity resulting in maximum lateral CG of helicopter.

Weight (Pounds)

CG (Inches)

68 136 204 272 340 371 408 476 544 612 680 748 816 884 952 1020 1088 1156 1174 1224 1292 1360 1399 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2108 2176 2244 2312 2380 2448 2469

0 0 0 0 0 0 -0.03 -0.06 -0.05 -0.04 -0.04 -0.03 -0.03 -0.03 -0.03 -0.03 -0.02 -0.02 -0.02 -0.03 -0.44 -0.55 -0.60 -0.59 -0.56 -0.54 -0.52 -0.50 -0.48 -0.46 -0.44 -0.43 -0.41 -0.40 -0.39 -0.38 -0.36 -0.35 -0.34 -0.34


Moment (In-Lb)

0 0 0 0

-12 29 -27 -24 -27 -22 -24 -27 -29 -31 -22 -23 -23 -37 -568 -748 -839 -843 -838 -845 -849 -850 -849 -845 -838 -848 -836 -843 -849 -853 -832 -833 -832 -839

at 15°C

(59F).


MANUFACTURER'S DATA

Table 1-2M.

Fuel loading with left and right auxiliary tanks (two 61.7 liter) lateral (Metric)

Lateral

Jet B or JP4 (0.779

Quantity (liters)

40 80 120 160 200 206.6 240 280 * 302.7 320 360 400 440 480 520 560 600 640 * 653.2 680 720 760 * 778.6 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1374.1

Weight (kg)

31.2 62 3 93.5 124.6 155.8 161 0 187.0 218.1 235.8 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 498.6 508.8 529.7 560.9 592.0 606.5 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 872.5 903.6 934.8 966.0 997.1 1028.3 1059.4 1070.4

Jet A. A1 or JP-5 (0.815 kg/liter)

kg/liter)

CG (mm)

0 0 0 0 0 0 -1 -1 -1 1 -1 1 -1 1 -1 -1 1 -1 -1 -8 -11 -14 -15 15 14 13 -13 -12 -12 11 -11 -11 10 -10 -10 -9 -9 -9 -9

Moment (kg-mm)

Quantity (liters)

Weight (kg)

0 0 0 0 0 0 -187 -218 -236 -249 -280 -312 -343 -374 -405 -436 -467 -499 509 -4238 -6170 -8288 -9098 -9348 9162 8912 -9317 -8974 -9348 -8912 -9254 -9598 9036 9348 -9660 -8974 -9255 -9535 -9634

40 80 120 160 200 206.6 240 280 302.7 320 360 400 440 480 520 560 600 640 653.2 680 720 760 778.6 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1374.1

32.6 65.2 97.8 130.4 163.0 168.4 195.6 228.2 246.7 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 532.3 554.2 586.8 619.4 634.5 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 912.8 945.4 978.0 1010.6 1043.2 1075.8 1108.1 1119.9

*

'

'

'

"

NOTE: All data above represents usable fuel (basic and auxiliary) * Quantity at which CG of fuel changes direction. 1 Quantity resulting in maximum lateral CG of helicopter.

Moment (kg-mm)

CG (mm)

0 0 0 0 0 0 1 1 1 -1 -1 -1 -1 -1 1 -1 -1 -1 -1 -8 -11 14 15 -15 -14 -13 13 -12 -12 11 -11 -11 -10 -10 -10 -9 -9 -9 -9


at 15°C

0 0 0 0 0 0 -196 -228 -247 -261 -293 -326 -359 -391 -424 456 -489 -522 -532 -4434 -6455 -8672 -9518 -9780 -9584 -9324 -9747 -9389 -9780 -9324 -9682 -10041 -9454 9780 -10106 -9389 -9682 -9973 -10079

(59°F).

1-5

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-25.2, 25.3 AND25.4

MANUFACTURER'SDATA

Table1-3. Fuelloadingwith left orrightauxiliarytanks (one16.3 gal.) longitudinal(English)

Longitudinal


*

*

Jet A, Al or JP-5 (6.8 Lb/U.S. Gallon)


Weight (Pounds)

CG (Inches)

Moment (In-Lb)

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 156.3 160 170 180 189.4 200 210 220 230 240 250 259.4 270 280 290 300 311.4 320 330 340 346.7

65 130 195 260 325 379 390 455 520 585 650 715 780 845 910 975 1016 1040 1105 1170 1231 1300 1365 1430 1495 1560 1625 1686 1755 1820 1885 1950 2024 2080 2145 2210 2254

139.4 139.6 139.8 139.9 139.9 139.9 140.9 145.8 149.0 151.5 153.2 154.8 156.0 157.0 158.2 159.2 159.7 158.0 154.4 150.9 148.1 149.6 150.8 151.9 152.9 153.9 154.9 155.5 154.1 152.6 151.2 150.0 148.6 149.3 150.1 150.9 151.4

9061 18148 27261 36374 45468 53022 54951 66339 77480 88628 99580 110682 121680 132665 143962 155220 162255 164320 170612 176553 182311 194480 205842 217217 228586 240084 251713 262173 270446 277732 285012 292500 300766 310544 321965 333489 341256

NOTE: All data above * Quantity at which Quantity resulting Quantity resulting Quantity resulting

Quantity (U.S. Gal)

*

'

10 20 30 40 50 58.3 60 70 80 90 100 110 120 130 140 150 156.3 160 170 180 189.4 200 210 220 230 240 250 259.4 270 280 290 300 311.4 320 330 340 346.7

Weight (Pounds)

CG (Inches)

Moment (In-Lb)

68 136 204 272 340 397 408 476 544 612 680 748 816 884 952 1020 1063 1088 1156 1224 1288 1360 1428 1496 1564 1632 1700 1764 1836 1904 1972 2040 2117 2176 2244 2312 2358

139.4 139.6 139.8 139.9 139.9 139.9 140.9 145.8 149.0 151.5 153.2 154.8 156.0 157.0 158.2 159.2 159.7 158.0 154.4 150.9 148.1 149.6 150.8 151.9 152.9 153.9 154.9 155.5 154.1 152.6 151.2 150.0 148.6 149.3 150.1 150.9 151.4

9479 18986 28519 38053 47566 55540 57487 69401 81056 92718 104176 115790 127296 138788 150606 162384 169761 171904 178486 184702 190753 203456 215342 227242 239136 251165 263330 274302 282928 290550 298166 306000

represents usable fuel (basic and auxiliary) based on nominal density CG of fuel changes direction. in maximum forward CG of helicopter (at any weight). in maximum aft CG of helicopter (weight empty 6630 lb or less). in maximum aft CG of helicopter (weight empty more than 6630 lb).

314586 324877 336824 348881 357001

at 15°C (59°F).


MANUFACTURER'S DATA

Fuel loading with left or right auxiliary tanks (one 61.7 liter) longitudinal (Metric)

Table 1-3M.

Longitudinal



Quantity

Weight

CG

Moment

Quantity

Weight

CG

Moment

(liters)

(kg)

(mm)

(kg-mm)

(liters)

(kg)

(mm)

(kg-mm)

40 80 120 160

31.2 62.3 93.5 124.6

3541 3547 3551 3552

110479 220978 332019 442579

40 80 120 160

32.6 65.2 97.8 130.4

3541 3547 3551 3552

115437 231264 347288 463181

200

155.8

3552

553402

200

163.0

3552

220.7 240 280 320 360 400 440 480 520 560 591.5 600 640 680 717.1 720 760 800 840 880 920 960 982.0 1000 1040 1080 1120 1160 1178.5 1200 1240 1280

171.9 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 460.8 467.4 498.6 529.7 558.6 560.9 592.0 623.2 654.4 685.5 716.7 747.8 765.0 779.0 810.2 841.3 872.5 903.6 918.1 934.8 966.0 997.1

3553 3637 3739 3815 3871 3917 3952 3983 4013 4036 4057 4046 3927 3838 3762 3764 3800 3833 3861 3889 3917 3940 3950 3934 3894 3856 3823 3790 3775 3787 3807 3830

610761 680119 815476 951080 1085428 1220537 1354746 1489244 1625666 1760503 1869466 1891100 1958002 2032989 2101453 2111228 2249600 2388726 2526638 2665910 2807314 2946332 3021750 3064586 3154919 3244053 3335568 3424644 3465828 3540088 3677562 3818893

179.9 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 482.1 489.0 521.6 554.2 584.4 586.8 619.4 652.0 684.6 717.2 749.8 782.4 800.3 815.0 847.6 880.2 912.8 945.4 960.5 978.0 1010.6 1043.2

3553 3637 3739 3815 3871 3917 3952 3983 4013 4036 4057 4046 3927 3838 3762 3764 3800 3833 3861 3889 3917 3940 3950 3934 3894 3856 3823 3790 3775 3787 3807 3830

639185 711397 853240 994952 1135751 1276942 1417187 1558150 1700709 1842030 1955880 1978494 2048323 2127020 2198513 2208715 2353720 2499116 2643241 2789191 2936967 3082656 3161185 3206210 3300554 3394051 3489634 3583066 3625888 3703686 3847354 3995456

1022.3

3845

3930744

1069.5

3845

4112228

1312.3

NOTE: All data above Quantity at which Quantity resulting l Quantity resulting Quantity resulting

A

220.7 240 280 320 360 400 440 480 520 560 591.5 600 640 680 717.1 720 760 800 840 880 920 960 982.0 1000 1040 1080 1120 1160 1178.5 1200 1240 1280

1312.3

represents usable fuel (basic and auxiliary) based on nominal CG of fuel changes direction. in maximum forward CG of helicopter (at any weight). in maximum aft CG of helicopter (weight empty 3007 kg or less). in maximum aft CG of helicopter (weight empty more than 3007 kg).

density

578976

at 15°C (59°F).

1-7


MANUFACTURER'S DATA

Table 1-4. Fuelloadingwith left or rightauxiliarytanks(one 16.3 gal.) lateral(English)(Sheet 1 of 2)

Lateral



Weight (Pounds)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 152.6 156.3 157.8 160 170 180 182.8 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 340 346.7

+

Quantity Quantity Quantity Quantity Quantity

65 130 195 260 325 355 390 455 520 585 650 715 780 845 910 975 992 1016 1026 1040 1105 1170 1188 1235 1300 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2254

at which at which at which resulting resulting

CG of fuel changes CG of fuel changes CG of fuel changes in maximum lateral in maximum lateral

Lateral 116.3 Left) CG Moment (In-Lb) (Inches) 0 0 0 0 0 0 -0.04 -0.43 0.84 1.20 -1.58 -1.82 -2.07 -2.27 -2.43 -2.52 -2.53 -2.48 -2.67 -2.69 -2.69 2.69 -2.68 -2.67 -2.54 -2.42 -2.30 -2.20 -2.11 2.03 -1.95 -1.88 -1.81 -1.75 -1.69 -1.64 -1.58 -1.54 -1.49 -1.46

0 0 0 0 0 0 -16 -196 437 -702 -1027 -1301 -1615 -1918 -2211 -2457 -2510 -2520 -2739 -2798 2972 3147 -3184 -3297 -3302 -3303 -3289 -3289 -3292 -3299 -3296 -3299 -3294 -3299 -3296 -3305 -3286 -3303 -3293 -3291

direction. direction for left auxiliary tank only. direction for right auxiliary tank only. CG of helicopter for left auxiliary tank. CG of helicopter for right auxiliary tank.

Lateral (16.3 Right) CG Moment (inches) (In-Lb) 0 0 0 0 0 0 -0.05 0.38 0.80 1.15 1.50 1.77 2.01 2.22 2.39 2.49 2.49

0 0 0 0 0 0 -20 173 416 673 975 1266 1568 1876 2175 2428 2470

2.13 1.83 1.57

2215 2022 1837

1.37 1.29 1.23 1.18 1.13 1.08 1.04 1.00 0.96 0.92 0.89 0.86 0.84 0.81 0.78 0.76 0.75

1692 1677 1679 1687 1689 1685 1690 1690 1685 1674 1678 1677 1693 1685 1673 1680 1691


MANUFACTURER'S DATA Table 1-4.

Fuel loading with left or right auxiliary tanks (one 16.3 gal.) lateral (English) (Sheet 2)

Lateral

Jet A, A1 or JP-5 (6.8 Lb/U.S. Gallon)


Weight (Pounds)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 152.6 156.3 157.8 160 170 180 182.8 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 340 346.7

+

+

Quantity Quantity Quantity Quantity Quantity

68 136 204 272 340 371 408 476 544 612 680 748 816 884 952 1020 1038 1063 1073 1088 1156 1224 1243 1292 1360 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2108 2176 2244 2312 2358

at which at which at which resulting resulting

CG of fuel changes CG of fuel changes CG of fuel changes in maximum lateral in maximum lateral

Lateral (16.3 Left) CG Moment (Inches) (In-Lb) 0 0 0 0 0 0 -0.04 -0.43 -0.84 1.20 -1.58 -1.82 -2.07 -2.27 2.43 2.52 2.53 2.48 -2.67 -2.69 2.69 -2.69 -2.68 -2.67 -2.54 -2.42 2.30 2.20 -2.11 -2.03 -1.95 -1.88 -1.81 -1.75 -1.69 -1.64 1.58 -1.54 -1.49 -1.46

0 0 0 0 0 0 -16 -205 -457 -734 -1074 -1361 -1689 -2007 -2313 -2570 -2626 -2636 -2865 -2927 -3110 -3293 -3331 3450 -3454 -3456 -3441 -3441 -3444 3451 -3448 -3452 3446 3451 -3448 -3457 -3438 -3456 -3445 -3443

Lateral 116.3 Right) CG Moment (Inches) (In-Lb) 0 0 0 0 0 0 -0.05 0.38 0.80 1.15 1.50 1.77 2.01 2.22 2.39 2.49 2.49

0 0 0 0 0 0 -20 181 435 704 1020 1324 1640 1962 2275 2540 2585

2.13 1.83 1.57

2317 2115 1922

1.37 1.29 1.23 1.18 1.13 1.08 1.04 1.00 0.96 0.92 0.89 0.86 0.84 0.81 0.78 0.76 0.75

1770 1754 1756 1765 1767 1763 1768 1768 1763 1752 1755 1754 1771 1763 1750 1757 1769

direction. direction for left auxiliary tank only. direction for right auxiliary tank only. CG of helicopter for left auxiliary tank. CG of helicopter for right auxiliary tank.

1-9


BHT-412-FMS-25.2, 25.3 AND 25.4

Table 1-4M.

Fuel loading with left or right auxiliary tanks (one 61.7 liter) lateral (Metric) (Sheet 1 of 2)

Lateral


Weight (kg)

Quantity (liters)

31.2 62.3 93.5 124.6 155.8 160.9 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 450.0 460.8 465.1 467.4 498.6 529.7 538.9 560.9 592.0 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 872.5 903.6 934.8 966.0 977.1 1022.3

40 80 120 160 200 206.6 240 280 320 360 400 440 480 520 560 577.6 591.5 597.1 600 640 680 691.8 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1312.3

+

* ** +

Quantity Quantity Quantity Quantity

at which at which resulting resulting

CG of fuel changes CG of fuel changes in maximum lateral in maximum lateral

Lateral 161.7 Left) Moment CG (kg-mm) (mm) 0 0 0 0 0 0 -4 -15 -26 -35 -43 -50 -56 -61 -64 -64 -63 -68 -68 -68 -68 -68 -68 -64 -61 -58 -55 -53 -51 -49 -47 -45 -43 -42 -41 -39 -38 -37

0 0 0 0 0 0 -748 -3272 -6482 -9814 -13399 -17140 -20938 -24711 -27917 -28800 -29030 -31627 -31783 -33905 -36020 36645 -38141 -37888 -38015 -37955 -37703 -37985 -38138 -38171 -38079 -37859 -37518 -37951 -38327 -37674 -37890 -37825

direction. direction for left auxiliary tank only. CG of helicopter for left auxiliary tank. CG of helicopter for right auxiliary tank.

Lateral (61.7 Right) CG Moment (kg-mm) (mm) 0 0 0 0 0 0 2 14 25 34 42 49 55 60 63 63

0 0 0 0 0 0 374 3053 6233 9534 13087 16797 20565 24306 27481 28350

55 47 40

25707 23434 21188

35 33 31 30 28 27 26 25 24 23 22 21 21 20 19 19

19632 19536 19319 19632 19194 19351 19443 19475 19445 19350 19195 18976 19631 19320 18945 19424


Table 1-4M.

BHT-412-FMS-25.2, 25.3 AND 25.4

Fuel loading with left or right auxiliary tanks (one 61.7 liter) lateral (Metric) (Sheet 2)

Lateral


Weight (kg)

Quantity (liters)

+

** +

40 80 120 160 200 206.6 240 280 320 360 400 440 480 520 560 577.6 591.5 597.1 600 640 680 691.8 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1312.3

Quantity Quantity Quantity Quantity

32.6 65.2 97.8 130.4 163.0 168.4 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 470.7 482.1 486.6 489.0 521.6 554.2 563.8 586.8 619.4 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 912.8 945.4 978.0 1010.6 1043.2 1069.5

at which at which resulting resulting

CG of fuel changes CG of fuel changes in maximum lateral in maximum lateral

Lateral (61.7 Left) CG Moment (mm) (kg-mm) 0 0 0 0 0 0 -4 -15 -26 -35 -43 -50 -56 -61 -64 -64 -63 -68 -68 -68 -68 -68 68 -64 -61 -58 -55 -53 -51 -49 -47 -45 -43 -42 -41 -39 -38 -37

0 0 0 0 0 0 -782 -3423 -6781 -10269 -14018 -17930 -21907 -25852 -29210 -30125 -30372 -33089 -33252 -35469 -37686 -38338 -39902 -39642 -39772 -39707 -39446 -39739 -39902 -39935 -39837 -39609 -39250 -39707 -40098 -39413 -39642 -39572

Lateral (61.7 Right) CG Moment (mm) (kg-mm) 0 0 0 0 0 0 2 14 25 34 42 49 55 60 63 63

0 0 0 0 0 0 391 3195 6520 9976 13692 17571 21516 25428 28753 29654

55 47 40

26895 24515 22168

35 33 31 30 28 27 26 25 24 23 22 21 21 20 19 19

20538 20440 20212 20538 20082 20245 20342 20375 20342 20245 20082 19853 20538 20212 19821 20321

direction. direction for left auxiliary tank only. CG of helicopter for left auxiliary tank. CG of helicopter for right auxiliary tank.

1-11/1-12



TWO-SPEEDINTERNALHOIST (412-899-223) OR (214-706-003)

CERTIFIED SEPTEMBER 19, 1988

This supplement shall be attached to the Model 412 & 412EP Flight Manual when the 412-899223 or 214-706-003 Internal Hoist has been installed. Information contained herein supplements of basic Flight Manual. For information Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

ANDBELLHELICOPTER TEXTRON INC.

A DIVISION OFTEXTRON CANADA LTD ALLRIGHTS RESERVED

REISSUE -

MAY 11,

1995


FAA APPROVED

NOTICEPAGE





FAA APPROVED

LOG OF REVISIONS


.... September 19, 1988

Revision..........1 Reissue ...........

........... 0 ...........

May 25, 1990 May 11, 1995

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP .............................. A - B .................................. i/ii ....................................... 1 - 34 ..................................

REVISION NO.

PAGE

35/36 ....................................

0 0 0 0

0

MANUFACTURER'S DATA

37 -38

.................................

0

NOTE



BHT-412-FMS-26

REVISIONS LOGOFAPPROVED

Original ........... Revision .......... Reissue...........0

.... September 19, 1988 May 25, 1990 1 ........... ........... May 11, 1995

APPROVED:

1 MANAGER


B

FAA APPROVED


GENERAL INFORMATION

The Two Speed Internal Hoist enables cargo and emergency rescue operations In areas where landing cannot be accomplished. The hoist can raise or lower loads up to 600 pounds (272 kilograms). The hoist contains 250 usable feet (76.2 meters) cable. Each of the four cabin mounting locations allows the hoist to be extended 90 degrees outboard. The hoist provides two extend/retrieve speeds (HIGH and LOW). With LOW speed selected, a

continuously variable speed range from zero to 125 feet/minute (38.1 meters/ minute) is available. With HIGH speed selected, a continuously variable speed range from zero to 250 feet/minute (76.2 meters/minute) is available. An electrically actuated cable cutting device allows either the pilot or hoist operator to sever the cable if necessary. A manually actuated cutting device is provided for use in the event of an electrical failure.

i/ii

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-26

1-3. TYPESOF OPERATION

Hoist operation with flight director in coupled mode is prohibited.

Hoist operations shall be conducted under appropriate operating rules for external loads.

Hoisting or lowering an empty litter in open position (except Stokes litter) is prohibited.

Passenger operations with hoist installed are approved if hoist is stowed and electrical system is deactivated.

Refer to appropriate Flight Manual Supplement(s) for additional limitations, procedures, and performance data.

Hoist operations are prohibited during instrument meteorological conditions.

1-4. FLIGHTCREW

A crewmember wearing an approved safety harness in passenger compartment is required during all phases of hoist operations. Crewmember shall wear protective gloves for guiding cable during operation. The hoist operator shall be familiar with hoist operating procedures and limitations.

1-5. CONFIGURATION

1-5-A. REQUIRED EQUIPMENT

Hoist cable antichafing guard shall be installed on standard or high skid landing gear (with or without floats) on same side of helicopter as hoist.

1-5-B. OPTIONALEQUIPMENT

Fixed passenger step shall not be installed concurrently with internal hoist.

1-6. WEIGHTAND CENTEROF GRAVITY

Actual weight change shall be determined after hoist is installed and ballast readjusted, if necessary. For maximum gross weight, including hoist load, refer to applicable Flight Manual or BHT-412-FM-19.1 when Increased Gross Weight and Takeoff Horsepower kit is installed.

Maximum hoist load is 600 pounds (272 kilograms). This is a structural limitation only and does not ensure that longitudinal or lateral CG will remain within approved limits. Maximum allowable hoist load varies with gross weight, center of gravity, and hoist location. Refer to appropriate Hoist Loading Schedule. NOTE

The center of gravity of hoist load in forward position is F.S. 82 (2083 mm) and B.L. 60 (1524 mm). The center of gravity of hoist load in aft position is F.S. 131 (3327 mm) and B.L. 64.4 (1636 mm).

Retractable passenger steps shall be stowed during hoist operations. 1

T N I R P N O R I O T F C U T O OD N R E P E E I R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-26

FAA APPROVED

For Longitudinal vs. Lateral CG limits with Internal hoist refer to Internal hoist CG envelope figure 1-1.

1-7. AIRSPEED

VNE with asymmetrical door configuration is 20 KIAS.

VNE with hinged panels locked open and cargo doors open is 20 KIAS. VNE with hinged panels removed and cargo doors removed or secured open is 60 KIAS.

1-23. HOIST SPEED

HIGH speed - Limited to hoist loads of 300 pounds (136 kilograms) or less.

LOW speed - Limits of basic hoist (600 pounds., 272 kilograms).

1-24. HOIST DUTYCYCLE LIMITATIONS

The hoist is approved for continuous operation with loads not to exceed 600 pounds (272 kilograms).

1-25. ALLOWABLE HOIST LOAD

Select hoist loading schedules (figures 1-2 through 1-5) appropriate for position in which hoist is installed. NOTE

Hoist loading schedules are based on most adverse loading combinations of pilot, copilot, and hoist operator, each weighing 170 or 200 pounds (77.1 or 90.7 kilograms), and on a weight empty CG of 0.3 inches (7.3 mm) to right

2

of centerline prior to adding hoist. If lateral CG is appreciably different or crewmember weights are out of this range, allowable hoist load shall be computed. For computation, assume hoist operator in forward position to be located at F.S. 87 (2210 mm) and B.L. 40 (1016mm), and in aft position F.S. 125 (3175mm) and B.L. 40 (1016mm).

1-25-A. LEFT HOIST INSTALLATIONS

Enter appropriate schedule, figures 1-2 through 1-5 at gross weight of helicopter prior to hoisting. Proceed vertically to intersect with diagonal line representing number of crewmembers on board, top of schedule, or right cutoff line. Proceed horizontally to left to read maximum allowable hoist load. Intersecting with right cutoff line gives maximum load which does not cause helicopter to exceed gross weight limitations.

Using Weight empty chart, Section 5 and left hoist loading schedules ensures that both longitudinal and lateral limits are not exceeded during first hoist operation. However, for subsequent hoisting, additional precautions must be taken to avoid exceeding forward longitudinal limits.

1-25-A-1. LEFT FORWARD HOIST LOCATION

To continue using maximum allowable hoist capability: (Refer to figure 1-2 through 1-5)

a. put hoisted load (people or cargo) along side of island, or

b. when hoisted load is put immediately forward of island, reduce maximum hoist load to 300 pounds.


WARNING

DO NOT PUT HOISTED LOAD IN FORWARDAREA OF PASSENGER COMPARTMENT UNLESS MAXIMUM HOIST LOADS ARE COMPUTED FOR THAT CONFIGURATION.

1-25-A-2. LEFT AFT HOIST

LOCATION

To continue using maximum allowable hoist capability: (Refer to figure 1-2 through 1-5) a. put hoisted load along island or immediately forward of island, or b. ensure empty weight CG Is within Area A. Refer to Weight empty chart, Section 5.

1-25-B. RIGHTHOIST INSTALLATIONS - NORMAL OPERATIONS

Right lateral limit for hoist operations varies with longitudinal center of gravity of the helicopter. The loading schedules have been modified to account for this variation.

BHT-412-FMS-26

Intersecting right cutoff line gives maximum load which does not cause helicopter to exceed gross weight limitations or forward longitudinal limits.

For multiple hoists during a single flight, after each hoist operation enter appropriate schedule at revised gross weight and proceed to new center of gravity to determine maximum allowable hoist load. EXAMPLE 1: NORMAL

Determine Hoist Load when hoist Is in R/H FWD POISTION and crew consist of Pilot, Copilot and Hoist Operator. GIVEN:



Enter gross weight at 9,500 lbs. Proceed up GW line to interpolated STA. 135.5 Proceed left to read hoist load of 210 lbs. Point .

EXAMPLE2: NORMAL

a. Starting with appropriate schedule for number of crewmembers on board, enter at gross weight of helicopter prior to hoisting. b. Proceed vertically to Intersect with diagonal line representing helicopter center of gravity prior to hoisting, top of schedule, or right cutoff line. c. Proceed horizontally to left to read maximum allowable hoist load.

Determine Hoist Load when hoist is in R/H FWD POSITIONand crew consist of Pilot, Copilot and Hoist Operator.

When helicopter center of gravity is between STA. lines, interpolate to determine CG.

Enter gross weight at 9,500 lbs. Proceed up GW line to STA. 138.5 Proceed left to read hoist load of 550 lbs. Point

GIVEN:



3


FAA APPROVED

1-25-C. RIGHTHOIST

INSTALLATION - PENALTY REGION

OPERATION

The dashed line on schedules represents longitudinal center of gravity prior to hoisting which will result In a gross weight center of gravity at Sta. 135.2 and B.L. 4.5 during hoist operations with maximum hoist loads derived using this line. This center of gravity is the corner of but not in Penalty Region shown in Limitations. Hoist loads derived for Normal Operations may be Increased when GW/CG combinations are forward of those represented by dashed line. Loads may be increased up to but not greater than those defined by dashed line. However, this procedure will result in operations within Penalty Region. Refer to Section 1, Internal Hoist CG Envelope, for Penalty Region. EXAMPLE3: PENALTYREGION

Determine Hoist Load when hoist Is In R/H FWD POSITIONand crew consist of Pilot, Copilot and Hoist Operator. GIVEN:

Gross Weight CG -

9,500 lbs.

STA. 135.5 before hoisting

From appropriate 11,600 lb. GW schedule obtain hoist load as previously determined In Example 1 the maximum hoist load for normal operations Is 210 lbs. Point (A).

To increase hoist load to maximum for condition without exceeding GW/CG limits, proceed up to dashed line and read left to find 435 lbs. Point (c). The Penalty Region is any load greater than Point (A) up to maximun load at Point (c). For GW vs. CG combinations aft of the CG represented by the dashed line (see Example 2), there Is no Penalty Region.

1-26. WEIGHT EMPTY CHART

The Weight empty chart for internal hoisting operations is shown In Section 5. Refer to the maintenance manual for additional Information. NOTE

Allowable hoist load must be computed when weight empty is not within specified guidelines, shown in Section 5. NOTE

Allowable hoist loads must be computed when AUX Fuel kits are Installed.


BHT-412-FMS-26

;

Figure 1-1. Internal hoist C.G. envelope

5


I

FAA APPROVED

Figure 1-2. Hoist loading schedules 11,600 lb. GW (English) (Sheet 1 of 6)

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V .. NLY O BHT-412-FMS-26

FAA APPROVED

o

m

0r

o

0

o


7


FAA APPROVED


8


w

Figure 1-2.


9


FAA APPROVED o

10


BHT-412-FMS-26

o

Figure 1-2. Hoist loading schedules 11,600lb. GW (English) (Sheet 6 of 6)



BHT-412-FMS-26

13


FAA APPROVED

Figure 1-3. Hoist loading schedules 5261 kg. GW (Metric) (Sheet 3 of 6)

14


BHT-412-FMS-26

15


Figure 1-3.



BHT-412-FMS-26


17


FAA APPROVED

I


18


BHT-412-FMS-26

o

oo

,

O

0

N

Figure 1-4. Hoist loading schedules11,900 lb. GW (English) (Sheet 2 of 6)

19


FAA APPROVED

,

w

Z


20


BHT-412-FMS-26

o

0

I

Figure 1-4.

Hoist loading schedules

11,900 lb. GW (English) (Sheet 4 of 6)

21


FAA APPROVED

I

0

0

l

0 0

0

I

I I

Figure 1-4.

22



BHT-412-FMS-26

o

0

]

o

LO

u

Mn

z

0

0

Figure 1-4. Hoist loading schedules11,900 lb. GW (English) (Sheet 6 of 6)

23


FAA APPROVED

0

eI.

-

o

o

0

Figure 1-5.

24


I R N P O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O L/H AFT POSITION

1 PILOTAND HOISTOPERATOR-272 KG HOISTLOADIS ALLOWEDATGROSSWEIGHTSFROM3498.2TO 5125.6KGS. 2 PILOT,COPILOT,AND HOISTOPERATOR- 272 KG HOISTLOADIS ALLOWEDATGROSSWEIGHTSFROM3830.8TO 5125.6KGS. 5125.6

250

100

0

-

3000

3250

3500 3750 4000 4250 4500 4750 5000 GROSSWEIGHTPRIORTOHOISTOPERATION - KILOGRAMS

5250

5500

412FS-26-1-5-2

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O R/H FWD POSITION - PILOT AND HOIST OPERATOR

5125.6

250

THOIST LOADING

PROHIBITED - KILOGRAMS GROSSWEIGHTPRIORTO HOISTOPERATION NOTE: THE DASHED LINE INDICATES A DIFFERENT LONGITUDINAL C G. FOR EACH GROSS WEIGHT. |

INOTE:

THE

DASHED

LINE

INDICATES

A DIFFERENT

LONGITUDINAL

C.G.

FOR

EACH

GROSS

L

IG


BHT-412-FMS-26


27


FAA APPROVED

0

0

Figure 1-5.



BHT-412-FMS-26

o

Figure 1-5.


29

T N I Section 2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-26

FAA APPROVED

2-2. FLIGHTPLANNING WARNING

HOIST LOAD CAN CAUSE LONGITUDINALOR LATERAL CG LIMITS TO BE EXCEEDED.GROSS WEIGHT AND CENTER OF GRAVITY SHALL BE COMPUTED TO ASSURE LOADING WITHIN APPROVEDLIMITS.

CAUTION

IF ADDITIONAL LOADS ARE CARRIED DURING HOISTING OPERATIONS, LOADS SHOULD BE PLACED ON SIDE OF HELICOPTER OPPOSITE HOIST POSITION.

Gross weight and CG - Compute with and without hoist load.


2-4-A. HOISTINSTALLATION CHECK NOTE

If pilot plans to operate hoist with other crewmember in passenger compartment, hoist shall be installed in forward right position.

Hoist - Installed In desired position; check roof and floor stud adapters and locking collars properly secured. Boom actuator - Installed position; all fittings secured.

In proper

AIRCRAFT POSITION switch (on hoist control box, figure 2-1) - Set in proper position. Hook - Rotates freely on cable.

Cable - Check proper routing through guide rollers, pulleys, and drums.

Gearbox oil levels - Check sight glasses. Hoist operators pendant connectors secured.

Installed;

Electrical power cables connectors secured.

Condition;

WARNING

ACTUATION OF CABLE CUT SWITCH ON PEDESTALCAN CUT CABLE REGARDLESS OF HOIST PWR SWITCH POSITION. ACTUATION OF CABLE CUT SWITCH ON HOIST CAN CUT CABLE, EVEN IF CABLE CUT CIRCUITBREAKERIS OUT.

CABLE CUT switches (pedestal and hoist) - Off; covers safetied. Safety vests, tether straps, hoisting slings, and litters - Condition; secured or stored.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-26

2-4-B. HOIST OPERATION CHECK Cargo doors and hinged panels open or removed.

Secured

Reduce hoist speed as cable approaches up limit. Check that hoist stops when hook reaches up limit without excess tension on cable.

HOIST PWR.CONT and CABLE CUT circuit

Hoist - Stowed for flight; hook restraint secured.

BATTERY switches external power).

HOIST PWR switch - OFF.

breakers -

In.

ON (or connect

NON ESNTLBUS switch -

MANUAL.

ICS - Check intercom between pilot and hoist operator using hoist pendant ICS trigger and HOT MIC switch (right ICS box only). HOIST PWR switch - ON. Check that blue HOIST POWERlight on hoist control box and amber CAUTION light on hoist pendant illuminates.

Hoist pendant CAUTION and OVER TEMP indicators - Pres to test.

HOIST UP/DOWN, BOOM IN/OUT, and SPEED HIGH/LOW switches (pilot and operator) - Actuate to check all hoist functions for proper operation. Check that pilot HOIST switch overrides pendant HOISTswitch. Hoist OVERTEMPwarning lights to test.

Press

CAUTION

NON ESNTLBUS switch - NORMAL. BATTERY switches - OFF.

NOTE

Ground crewmember should be instructed to discharge helicopter static electricity before attaching load to hoist when possible.

2-6. SYSTEMSCHECK

Cargo doors and hinged panels - Secured open or removed. CABLE CUT switches (pedestal and hoist) - Off; covers safetied.

HOIST PWR, CONT, and CABLE CUT circuit breakers - In.

2-6-A. BEFORE TAKEOFF

Safety vests and straps - On and secured to helicopter. Gloves - On.

MAINTAIN TENSION ON HOIST CABLE WHILE REELING IN AND OUT TO PREVENTSLACK.

HOIST and BOOM switches (pilot and operator) - Actuate to check all hoist functions for proper operation. Check that pilot HOIST switch overrides operator pendant HOISTswitch. Hoist cable - Check for corrosion, kinks, flat spots, fraying, or broken strands.

Up limit switch actuator - Raise while hoist is reeling in and check hoist motor stops; then release and check hoist resumes operation.

STEP switch (If installed) - STOW.


Maximum hoist load shall be determined prior to each hoist operation. NOTE The Height-Velocity Diagram Is not a limitation for internal hoist operations under an appropriate operating certificate.

HOIST PWR switch - ON.

31


FAA APPROVED

WARNING

CAUTION

HOIST OPERATOR SHALL BE SECUREDTO HELICOPTERWITH AN APPROVED SAFETY HARNESS DURING HOIST OPERATIONS.

USE CARE TO PREVENTCABLE, AND LOAD FROM HOOK, FOULING ON FUSELAGE OR LANDINGGEAR.

Establish hover over hoist operation area.

Maintain zero ground speed until load is clear of obstructions.

Hoist hook restraint - Removed.

As desired (refer to

BOOMswitch - IN to swing hoist boom and load into cabin, if possible.

BOOM switch (or pilots HOIST switch) OUT.

Takeoff into wind, if possible, allowing adequate hoist load clearance over obstacles if load is not internal.

SPEED switch limitations).

NOTE

Each hoist operation performed Is defined as reeling hoist cable out and then in while hovering with any weight on hoist, regardless of whether the hoist was used for training or an actual rescue.

The pilot must record each operation In the penalty CG region. For each hoist operation performed within penalty CG region, four (4) additional hours of usage must be logged against the main rotor yoke, mast and lower cone seat. HOIST switch - DOWN.

Discharge static electricity when possible, and connect hook to load, observing allowable hoist load. NOTE

As hook nears the up or down limits, hoist speed automatically slows.

HOISTswitch - UP.

CAUTION

AIRSPEED WITH EXTERNAL BY LOAD IS LIMITED CONTROLLABILITY. CAUTION SHOULD BE EXERCISED WHEN CARRYING AN EXTERNAL LOAD. HANDLING CHARACTERISTICS MAY BE AFFECTEDBY THE SIZE, WEIGHT, AND SHAPE OF LOAD.

Airspeed - As required for adequate controllability, not to exceed limits for hoist operations (20 or 60 KIAS, as applicable).

2-13. LITTER HOISTING

of a When emergency transportation patient by litter is essential, every effort should be made to land the helicopter for litter loading. Litter hoisting can be hazardous and should be accomplished only when a landing is not feasible and the condition of the patient precludes the use of the personnel hoisting sling.

In addition to all other procedures contained herein, the following shall apply to litter hoisting operations. 32


2-13-A. EMPTY LITTER WARNING

HOISTING OR LOWERING AN EMPTY LITTER IN OPEN POSITION IS PROHIBITED. AN EMPTY LITTER CAN OSCILLATE UNCONTROLLABLY IN ROTOR WASH AND FLY UPWARD, STRIKING FUSELAGE OR TAIL ROTOR.

Prior to hoisting or lowering an empty litter, litter shall be closed and secured with straps. Litter should be suspended in a near-vertical position and sling straps should be drawn tight.

2-13-B. LOADED LITTER

BHT-412-FMS-26

SAFETY STRAPS.

HOIST HOOK CATCH SHALL BE SECURED WITH SAFETY PIN PRIOR TO HOISTING LITTER PATIENT.

Litter sling straps should be adjusted so that litter is 24 to 28 inches (61 to 71 centimeters) below hoist hook. NOTE

If litter is suspended too far below hook, litter cannot be loaded in helicopter with hoist hook at up limit.

CAUTION

.................

A LOADED LITTER CAN ROTATE CABLE DURING ABOUT HOISTING. HOIST OPERATOR MAY HAVE TO GRASP LITTER SLING STRAPS TO CONTROL ROTATION AS LITTER APPROACHES LANDING GEAR.

WARNING

LITTER PATIENT SHALL BE SECURED TO LITTER WITH

33


FAA APPROVED

AMBER-

ON SIDE OF HOIST

ATTACHED TO HOIST

TOP OFHOIST ON CONTROL BOX

ON

EXT

L

ARMED

L

OFF

IDLE STOP

ENG

ENG

START

ENG

NSRCH

UP

H

EXT

ENG

LT

O

HOIST PWR ON

OFF

PILOT COLLECTIVE

ON PEDESTAL

412FS--26-2-1

Figure 2-1.

34

Internal hoist controls

T Section3 N I R P N O R I O T F C U T O OD N R E P E E Section4 R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-26

FAA APPROVED

3-15. HOISTLOAD JETTISON

To jettison hoist load in an emergency, actuate CABLE CUT switch (located on pedestal or hoist).

In the event of failure of CABLE CUT switch, sever cable with manual cable cutter (stowed In pouch on hoist).

3-15-A. HOISTOVERTEMP WARNINGLIGHT

In the event that the OVERTEMPwarning light (located on pendant) illuminates, continue present operation until hoist cable Is reeled in. Leave HOIST PWR switch ON (for cooling fan operation). When OVERTEMP light extinguishes, hoisting may be resumed as desired.


35/36

T N I Section 5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

5-11. WEIGHTEMPTY CHART

BHT-412-FMS-26

NOTE

Allowable hoist load must be computed when AUX Fuel kits are installed.

The Weight empty chart for internal hoisting operations is shown in figure 5-1. Refer to the maintenance manual for additional information. NOTE

Allowable hoist load must be computed when weight empty is not within specified guidelines.

37


MANUFACTURER'S DATA

412 Weight Empty Chart for Internal Hoist Operations

8000 (3628.7)

Fwd Lin for Lef Aft Rit Fwd & Right Aft Positions

7500 (3401.9)

/

AREA

FwdLine

for L Fwd Positions

Line All Positions

7000

(3175.1)

(2948.3)

6000 (2721.5)

140 (3556)

141 (3581)

142 (3607)

143 (3632)

144 (3658)

145 (3683)

FUSELAGE STATION - INCHES (MILLIMETERS)

412FS-26-5-1

Figure 5-1.

38

Weight empty chart

T N I R P N O R I O T ROTORCRAFT F C MANUAL FLIGHT U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-27

MODELS

SUPPLEMENTFOR LITTERKIT OPERATION (205-706-047)

CERTIFIED 14 OCTOBER 1988

This supplement shall be attached to the Models 412 and 412 EP Flight Manual when the Litter Kit 205-706-047 has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter

T


COPYRIGHT NOTICE COPYRIGHT 1994 INC TEXTRON BELLHELICOPTER ANDBELL HELICOPTER TEXTRON, A DIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

POST OFFICE BOX 482 - FORT WORTH, TEXAS 76101

REISSUE -

23 JUNE 1994



BHT-412-FMS-27


LOG OF PAGES

Original .... Reissue.....

0 ............. 0 .............

14Oct 88 23Jun 94

FLIGHTMANUAL

Page

Title....................... A ......................... i/ii ......................... 1-1/1-2 ..................... 2-1/2-2 ..................... 3-1/3-2 ..................... 4-1/4-2....................

MANUFACTURER'SDATA

Revision No. 0 0 0 0 0 0 0

Page

1-1/1-2 .....................

APPROVED:

MANAGER


NOTE: Revisedtext is indicated by a black vertical line. Insertlatest revision pages;dispose of supersededpages.

Revision No.

0

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAAAPPROVED SUPPLEMENT


Introduction

BHT-412-FMS-27

INTRODUCTION

The Litter Kit provides three litters and the provisions for installing up to three litters in the helicopter. A cabin attendant seat is also includedin the kit.

i/ii


Section 1


BHT-412-FMS-27


MINIMUM FLIGHTCREW

Actual weight change shall be determined after the litter(s) are installed and ballast readjusted if necessary, to return empty weight CG within allowable limits.

The minimum flight crew for litter operations shall consist of a pilot and a second crewmember or cabin attendant, both of whom shall be trained in and capable of assisting in litter patient emergencyevacuation procedures.

1-1/1-2



Section 2

BHT-412-FMS-27

LITTERLOADING

Secure patients to litters, then load litters aboard the helicopter in sequence from top to bottom. When only two patients are carried, they should occupy the top and center litter positions. When only one patient is carried, the center litter should be used.

LITTERUNLOADING

opened. Refer to Section 3. Emergency Procedures for unloading procedures when cabin doors cannot be opened.

Open cabin door and unload litters and patients from the helicopter in sequence from bottom to top.

Litters to be handled by one person inside cabin and one person outside cabin.

NOTE

Normal unloading procedures apply when either passenger door can be

2-1/2-2



Section 3

BHT-412-FMS-27

IRNAIW

UNLOADINGTHROUGH EMERGENCYEXITS NOTE

In the event that cabin doors can not be opened,litter patients shall be unloaded through emergency pop-out windows. After all litter patients have been removed,ambulatory patients may then exit.

Removeemergency pop-out window by pushing at comers as marked.

Unstrap patient on center litter and remove patient through window opening.

Disconnect top litter at end near open window and lower end to rest on center litter. Remove patient retention straps and slide patient down litter and out through window opening. Raisetop and center litter ends near open window and engage center litter in brackets for top litter. Disconnect bottom litter. Raise bottom litter at end near open window and rest handles on the lower surface of the window opening. Unstrap patient and slide patient up litter and through window opening.

3-1/3-2

T N I Section4 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT


Section 4

BHT-412-FMS-27

No changefrom basic Flight Manual.

4-1/4-2

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 412 ROTORCRAFT MANUFACTURER'SDATA

Section 1

BHT-412-FMS-27

TABLEOF MOMENTS

TABLEOF MOMENTS

(IN-LB)

(kg * mm) 100

LITTERPATIENT

Weight (Pounds)

Loaded Laterally FS. 117

100 110 120 130 140 150 160 170 180 190 200 210 220

11700 12870 14040 15210 16380 17550 18720 19890 21060 22230 23400 24570 25740

LITTERPATIENT

Weight (K.G.)

50 55 60 65 70 75 77.1 80 85 90 95 100 105 110

Loaded Laterally 2972 mm 1486.0 1634.6 1783.2 1931.8 2080.4 2229.0 2291.4 2377.6 2526.2 2674.8 2823.4 2972.0 3120.6 3269.2

1-1/1-2

BHT-412-FMS-28.2 AND 28.3

T N I R P N O R I FLIGHTMANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODEL

33108-

33213

33214 - 33999 36020 AND SUB

SUPPLEMENTFOR DUAL BATTERYINSTALLATION (412-899-225) CERTIFIED APRIL5, 1989

Manual (BHT-412-FM-2 or -3) when the 412-899-225 Dual Battery Installation has been installed. The Information contained hereinsupplements the information of the basic Flight Manual. For Limitations, Procedures, and PerformanceData not contained in this supplement, consultthe basic Flight Manual.

ALL RIGHTSRESERVED

REISSUED

-

8 OCTOBER 1991


PROPRIETARY RIGHTSNOTICE

These data are proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation and/or maintenance is forbidden without prior written authorization from Bell Helicopter Textron, Inc.

Additionalcopiesof this publicationmay beobtainedby contacting: CommercialPublicationDistributionCenter BellHelicopterTextron P.O.Box482 FortWorth, Texas76101


BHT-412-FMS-28.2 AND 28.3

LOG OF REVISIONS

Original ........ Reissued .......

.......... 0 .......

April 5, 1989 October 8, 1991

LOG OF PAGES

FLIGHT MANUAL

MANUFACTURER'SDATA

REVISION NO.

PAGE

Title .......................... PN ........ A/B ........................... i/i ............................ 1-1/1-2 ........................ 2-1 - 2-2 ...................... 3-1 - 3-2 ...................... 4-1/4-2 ........................ 6-1 - 6-4 ......................

PAGE

REVISION NO.

0 1-1/1-2 ........................ 0..................02-1 - 2-4 ...................... 0 0 0 0 0 0 0

0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT. WORTH,TEXAS 76193-0170

NOTE Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. A/B


FAA APPROVED

INTRODUCTION

The dual battery system consists of two 25.2 volt, 40 ampere-hour batteries located in the nose compartment. Both batteries are connected to BAT BUS NO. 1 in parallel. BATT 1 HOT and BATT 2 HOT warnings are located on the caution segment panel. Battery switches are located on the instrument panel and are labeledBATT 1 and BATT 2. There are two circuit breakers, labeledBAT 2 CB1 and BAT 2 CB2, located on the nose compartment floor, on the right side of the helicopter.

i/ii


BHT-412-FMS-28.2 AND 28.3


Weight change shall be determined after the dual batteries have been installed, and ballast shall be readjusted (if necessary) to return empty weight CGto within allowable limits.

ELECTRICALLIMITATIONS BATTERY LIMITATIONS

BATTERYSHALL NOT BE USED FOR ENGINESTART AFTERILLUMINATION OF BATT 1 HOT OR BATT 2 HOT SEGMENTLIGHT.BATTERYSHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'SINSTRUCTIONS.

Maximum battery case temperature is 54.5 °C (130 °F), as indicated by illumination of BATT 1 HOTor BATT 2 HOT segment lights.

1-1/1-2


BHT-412-FMS-28.2 AND 28.3

BEFOREEXTERIORCHECK Flight planning - Completed.

Press-to-drainvalves - Press.

Grossweight and CG - Compute (refer to Weight and Balance section in Manufacturer's Data BHT-412-MD).

Fuel filters follows:

Drain before first flight of day as

BOOSTPUMPswitches - ON.

Publications- Checked.

Portable fire extinguishers secured.

Forward and middle fuel sumps - Drain samples as follows:

-

Installed and

Aft fuel sumps - Drainsamplesas follows: FUELTRANS switches - OFF

BOOSTPUMP switches - OFF FUELswitches - OFF. BATT 1 switch - ON.

FUELswitches - ON.

Fuel filter (left and right) - Drain samples. FUELswitches - OFF.

BOOSTPUMPswitches - OFF BATT 1 switch - OFF.

BATTERYBUS 1 switch - AUTO TRIPOFF

EXTERIORCHECK

BATTERYBUS 1 switch - ON.

Aft fuel sump drain buttons (left and right) Depress. NOTE

If aft sumps fail to drain, the sump valves may be operated manually.

1. FUSELAGE- FRONT

Nose compartment - Condition: connected: door secured.

batteries

NOTE

Ballastis not allowed beneath batteries.

2-1


PRESTARTCHECK

NOTE

BATT 1 switch and BATT 2 switch will be ON during GROUNDPOWERSTART. NOTE

Both batteries, or either single battery may be used for engine start when properly charged. This procedure shows a dual battery start.

BATT 1 switch - ON.

FAA APPROVED

GEN 1 switch increases.

ON: check ammeter load

BATT 2 switch - ON when ammeter load drops below 200 amps. NOTE

Before attempting generator assisted start on second engine, it is recommended that the battery be chargeduntil ammeter loaddrops below 150 amps.

AMPS 1 indicator - Check at or below 150 amps.

BATT2 switch - ON.

BATTERYBUS 1 and BUS 2 switches - ON; check BATTERYcaution light illuminates.

ENGINESTARTING

Lighting and miscellaneousswitches - OFF. BATT 1 and BATT2 switches - OFF.

ENGINE11START ENGINE START.

N1 RPM - Check 71% minimum. BATT 2 switch - OFF. NOTE

After start, in order to avoid excessive generator drive loads, charge only one battery at a time.

2-2

ENGINESHUTDOWN

NOTE

BATTERYBUS 2 switch will trip OFF automatically when BATT 1 and BATT 2 switches are turned OFF.

Collective down lock - Securedas desired.

FAA APPROVED

BHT-412-FMS-28.2 AND 28.3

T N Section 3 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 3-1.

Warning lights

PANELWORDING

FAULTCONDITION

CORRECTIVE ACTION

BATT 1 HOT or BATT 2 HOT

Battery overheating.

Affected battery switch OFF. Land as soon as practical.

BATTERY SHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATION OF BATT 1 HOT OR BATT 2 HOT LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURERS INSTRUCTIONS PRIOR TO RETURN TO SERVICE.

3-1

BHT-412-FMS-28.2 AND 28.3

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 3-2.

Caution lights

PANELWORDING

FAULTCONDITION

CORRECTIVEACTION

BATTERY

BATTERYBUS 1 and BUS 2 switches/relays in the same position.

Turn BATT BUS 1 switch ON and BATT BUS 2 switch OFF If light remains on, reverse switches.

3-2


FAA APPROVED

No changs from basic manual

4-1/4-2

FAA APPROVED

BHT-412-FMS-28.2 AND 28.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O EMERGENCYPROCEDURES

Table6-1. Warninglights - Takeoff prior to CDP

PANELWORDING

FAULTCONDITION

CORRECTIVEACTION



Land immediately. Affected battery switch - OFF. WARNING

BATTERYSHALLNOTBEUSED FOR ENGINE START AFTER ILLUMINATION OF BATT1 HOT OR BATT 2 HOT LIGHT. BATTERYSHALLBE REMOVED AND SERVICED IN WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURN TO SERVICE.

6-1

BHT-412-FMS-28.2 AND 28.3

FAA APPROVED


Table6-2. Warning lights - Takeoff after CDP

PANELWORDING

FAULTCONDITION

CORRECTIVE ACTION



Accelerate to VTOSS.Affected battery switch - OFF.Landas

soon as practcal.

WARNING

BATTERYSHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATION OF BATT 1 HOT OR BATT 2 HOT LIGHT. BATTERY SHALL BE REMOVED IN SERVICED AND WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURNTO SERVICE.

FAA APPROVED

BHT-412-FMS-28.2 AND 28.3


Table6-3. Warninglights - Landing prior to LDP

PANELWORDING

FAULTCONDITION

CORRECTIVE ACTION

BATT 1 HOTor BATT 2


Affected battery switch OFF. Land as soon as practical.

BATTERYSHALL NOT BE USED FOR ENGINE START AFTER ILLUMINATIONOF BATT 1 HOT OR BATT 2 HOT LIGHT. BATTERY SHALL BE REMOVED AND SERVICED IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURNTO SERVICE.

6-3

BHT-412-FMS-28.2 AND 28.3

FAA APPROVED


Table6-4. Warninglights - Landingafter LDP

PANELWORDING

FAULTCONDITION

CORRECTIVEACTION



Land immediately. Affected battery switch - OFF. WARNING

BATTERYSHALLNOTBEUSED FOR ENGINE START AFTER ILLUMINATION OFBATT1 HOT OR BATT 2 HOT LIGHT. BATTERYSHALLBEREMOVED AND SERVICED IN WITH ACCORDANCE MANUFACTURER'S INSTRUCTIONS PRIOR TO RETURN TO SERVICE.


BHT-412-FMS-28.2 AND 28.3

MANUFACTURER'SDATA

MANUFACTURER'SDATA

BHT-412-FMS-28.2 AND 28.3

T Section 2 N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

INSTRUMENTPANELAND CONSOLES Incorporation of Dual Battery Installation will alter the instrument panel and dc power system as shown In figures 2-1 thru 2-3.

412075-3S4

Figure 2-1.

Instrument panel

2-1

MANUFACTURER'SDATA

BHT-412-FMS-28.2 AND 28.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Figure 2-2.

Instrument panel

MANUFACTURER'SDATA

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O INV NO. 1 SW

FROM

Figure 2-3.

DC electrical system (Sheet 1 of 2)

2-3

MANUFACTURER'SDATA

BHT-412-FMS-28.2 AND 28.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O DUAL

STARTER

Figure 2-3.

DC electrical system (Sheet 2)

BHT-412-FMS-29.2 AND 29.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 36001 - 36019

AND

33214 - 33999

36020 AND SUB

SUPPLEMENTFOR REMOVALOF UPPERAFT CENTER FUEL CELL (412-899-227) CERTIFIED 23 MAY 1989

This supplement shall be attached to the Model 412 Flight Manual (BHT-412-FM-2or -3) when the upper aft center fuel cell has been removed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

ALL RIGHTS RESERVED

REISSUED -

8 OCTOBER1991

BHT-412-FMS-29.2 AND 29.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O PROPRIETARY RIGHTS NOTICE

These data are proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation and/or maintenance is forbidden without prior written authorization from Bell Helicopter Textron, Inc.

Additionalcopiesof this publicationmaybeobtainedby contacting: CommercialPublicationDistributionCenter BellHelicopterTextron P.O.Box482 FortWorth, Texas76101


BHT-412-FMS-29.2 AND 29.3

LOG OF REVISIONS

Original ........ Reissued .......

0 ......... 0 .......

May 23, 1989 October 8, 1991

LOG OF PAGES

FLIGHTMANUAL

MANUFACTURER'SDATA

REVISION NO.

PAGE

Title ........... PN ............ A/B ............ 1-2 ..........

...............

0 ............... 0 ............... 0 ............... 0

REVISION NO.

PAGE

3-16 .................. 17/18 ................... 19/20 ...................

...... ...... ......

.0 .0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT.WORTH, TEXAS 76193-0170

NOTE


A/B


FAA APPROVED

OPTIONAL EQUIPMENT

Auxiliary fuel kits shall not be installed when the aft upper center fuel cell is removed.

WEIGHT CG LIMITATIONS

Actual weight change shall be determined after the upper aft center fuel cell is removed and balance readjusted, if necessary, to retain empty weight CG within limits.

41289941

Figure 1. Placards and decals

1

T Section 2 N I R P N O R I O T F C U T OSectionO3 D N R E P E E R R C S N G O Section 4 N I R W O E I F V NLY O BHT-412-FMS-29.2 AND 29.3

FAA APPROVED

No change to basic manual.




MANUFACTURER'SDATA

MANUFACTURER'SDATA

WEIGHT GRAVITY

EMPTY

CENTER

OF

With the upper aft center fuel cell removed, the total net weight and C.G. changes are depicted in Table 1.

Table 1.

Weight and center of gravity changes

WEIGHTCHANGE (LBS) -28.9

LONGITUDINAL

ARM(IN) MOMENT

190.2 -5500.9

LATERAL 0.1

-2.1

VERTICAL 35.3 -1020.6

FUELLOADING

Due to the fuel flow sequencing between the tanks, the fuel loading CG will vary between fuselage station 139.4 and 153.9.

Critical fuel CG's are shown on fuel tables 1, 2, 2M and 3, 3M. With normal crew and passenger loading, gross weight CG will remain within limits at any fuel quantity. Figure 2 depicts fuel tank locations by station number.

3

MANUFACTURER'SDATA


NOTE

Station O (reference datum) is located 20 inches (508 millimeters) aft of the most forward point of the cabin nose.

FUSELAGE STATIONS

0

REF DATUM

PILOT

SEAT

412061-97

Figure 2.

4

Internal fuel tank station location


MANUFACTURER'SDATA

Table2.

Fuel loading table (English)

USABLE FUEL LOADING TABLE, UPPERAFT CENTERTANK REMOVED Horizontal (English)

Jet B. JP-4 (6.5 Lbs/Gal)

U.S.

Weight

70 80 90 100 110

55 520 585 650 715

CG

Moment

144.7 147.8 150.1 152.0 153.6

65839 76856 87809 98800 109824

816 884 952 987 1020 1088 1156 1224 1292 1306 1360 1428 1496 1564 1632 1660 1700 1768 1816

150.0 145.5 141.8 140.1 141.0 142.8 144.3 145.7 146.9 147.2 145.8 144.4 143.0 141.7 140.6 140.1 140.8 141.9 142.6

122400 128622 134994 138279 143820 155366 166811 178337 189795 192243 198288 206203 213928 221619 229459 232566 239360 250879 258962

Most forward C.G. condition at weight empties under 6260 pounds has no fuel. Most critical fuel amount for most forward C.G. condition at weight empties 6260 pounds or greater.

This table is invalid with auxiliary fuel tank(s) installed.

5

MANUFACTURER'SDATA


Table2M.

Fuel loading table (Metric)

USABLEFUELLOADINGTABLE,UPPERAFTCENTERTANKREMOVED Horizontal(Metric)

Jet B, JP-4 (0.779 kg/L.)

Jet A, A-1, JP-5 (0.815 kg/L.)

Weight

Most forward C.G. condition at weight empties under 2839 kilograms has no fuel.

6


MANUFACTURER'SDATA

Table 3.

10 20 30 40 50 58.3 60 70 80 90 100 110 112 120 130 140 145.2 150 160 170 180 190 200 210 220 230 240 250 260 267

Fuel loading table - lateral (English)

0

-0.75 -0.71 -0.67 -0.63 -0.60 -0.57 -0.54 -0.52 -0.50 -0.48 -0.46 -0.45


7


MANUFACTURER'SDATA

Table3M.

Fuel loading table - lateral (Metric)

USABLE FUEL LOADING TABLE, UPPERAFT CENTER TANK REMOVED Lateral (Metric)

Jet A, A-1, JP-5 10.815 kg/L.)

Jet B. JP-4 (0.779 kg/L.)

Liter

Weight (kg)

40 80 120 160 200 240 280 320 360 400 424.1 440 480 520 '549.6 560 600 640 680 720 760 800 840 880 920 960 1000 1010.7

31.2 62.3 93.5 124.6 155.8 187.0 213.1 249.3 280.4 311.6 330.4 342.8 373.9 405.1 428.1 436.2 467.4 498.6 529.7 560.9 592.0 623.2 654.4 685.5 716.7 747.8 779.0 787.3

CG (mm) 0 0 0 0 0 0 0 0 0 0 0 -10 -15 -19 -21 -21 -19 -18 -17 -16 -15 -14 -14 -13 -12 -12 -11 -11

Moment (kg mm)

Liter

Weight (kg)

0 0 0 0 0 0 0 0 0 0 0 -3428 -5609 -7697 -8990 -9160 -8881 -8975 -9005 -8974 -8880 -8725 -9162 -8912 -8600 -8974 -8569 -8660

40 80 120 160 200 240 280 320 360 400 424.1 440 480 520 *549.6 560 600 640 680 720 760 800 840 880 920 960 1000 1010.7

32.6 65.2 97.8 130.4 163.0 195.6 228.2 260.8 293.4 326.0 345.6 358.6 391.2 423.8 447.9 456.4 489.0 521.6 554.2 586.8 619.4 652.0 684.6 717.2 749.8 782.4 815.0 823.7

Most critical fuel amount for most lateral C.G. condition. Weights given are nominal weights at 15°C.

NOTE

This table Is invalid with auxiliary fuel tank(s) Installed.

CG (mm)

0 0 0 0 0 0 0 0 0 0 0 -10 -15 -19 -21 -21 -19 -18 -17 -16 -15 -14 -14 -13 -12 -12 -11 -11 -11

Moment (kgmm) 0 0 0 0 0 0 0 0 0 0 0 -3586 -5868 -8052 -9406 -9584 -9291 -9389 -9421 -9389 -9291 -9128 -9584 -9324 -8998 -9389 -8965 -9061


MANUFACTURER'SDATA

SAMPLE

(ENGLISH)

LOADING

PROBLEM

The helicopter is chartered to transport nine passengersand 180 pounds of baggageon a trip that will require approximately 220 U.S. gallons of fuel one way. The helicopter will be refueled and the 190-pound pilot will return alone. Determine extreme CG conditions for both flights.

OUTBOUNDFLIGHT

LONGITUDINAL

WEIGHT

7265

CG

142.3

25 190 850 680 180

LATERAL

MOMENT

1034116

4146 8930 99450 59160 46980

CG

-0.2

0 +22.0 0 0 0

-1631

0

+4180

0 0 0


9190

136.3

1252782

+Takeoff Fuel (267 U.S. Gallons)

1816

142.6

258962

11006

137.4

1511744

+0.2

+1732


9190

136.3

1252782

+0.3

+2549

+Critical Forward Fuel (244.1 U.S. Gallons)

1660

140.1

232566

10850

136.9

1485348

+0.2

+1736

9190

136.3

1252782

+0.3

+2549

320

139.9

44768

0

9510

136.4

1297550

+0.3

Takeoff Condition


Basic Operating Weight + Payload +Landing Fuel (47 U.S. Gallons) Landing Condition

+0.3

MOMENT

-0.45

+2549 -817

-0.49

-813

0

+2549

9

MANUFACTURER'SDATA

BHT-412-FMS-29.2 AND 29.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O RETURN FLIGHT

LATERAL

LONGITUDINAL

WEIGHT (lbs.)

CG (in.)

MOMENT (lbs in

CG (in.)

(

( 100)

Weight Empty +Oil +Pilot

7265

142.3

25 190

1034116

4146 8930

MOMENT (lbs in) 100 )

-0.2

-1631

0 +22.0

+4180

0


7480

140.0

1047192

+0.3

+2549

+Takeoff Fuel (267 U.S. Gallons)

1816

142.6

258962

-0.45

-817

Takeoff Condition

9296

140.5

1306154

+0.2

+1732


7480

140.0

1047192

+0.3

+2549

+Critical Forward Fuel (244.1 U.S. Gallons)

1660

140.1

232566

-0.49

-813


9140

140.0

1279758

+0.2

+1736


7480

140.0

1047192

+0.3

+2549

+Critical Aft Fuel (112 U.S. Gallons)

762

153.9

117272

0

Most Aft Condition

8242

141.3

1164464

+0.3

+2549


7480

140.0

1047192

+0.3

+2549

320

139.9

44768

0

7800

140.0

1091960

+0.3

+Landing Fuel (47U.S. Gallons) Landing Condition

10

0

0

+2549

MANUFACTURER'SDATA

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SAMPLE

(METRIC)

LOADING

BHT-412-FMS-29.2 AND 29.3

PROBLEM

The helicopter is chartered to transport nine passengers and 80 kg of baggage for a trip that will require approximately 830 liters of fuel one way. The helicopter will be refueled and the 90 kg pilot will return alone. Determine extreme CG conditions for both flights.

OUTBOUNDFLIGHT

LONGITUDINAL

WEIGHT (kg)

CG (mm)

MOMENT (kg-mm

Weight Empty

+Oil +Pilot +Passengers, (5 man seat) +Passengers, (4 man seat) +Baggage

Basic Operating Weight + Payload +Takeoff Fuel (1010.7 Liters)

Takeoff Condition


LATERAL

CG (mm) -6

11.3 90.0 375.0 300.0 80.0

485.3 1074.6 11145.0 6630.0 5296.0

4151.6

3462

143723.0

823.7

3622

29834.4

0 +559 0 0 0

+7

-11

MOMENT

-198

0 +503 0 0 0

+305 -91

MANUFACTURER'SDATA

BHT-412-FMS-29.2 AND 29.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O RETURNFLIGHT

LATERAL

LONGITUDINAL

WEIGHT

Weight Empty +Oil +Pilot

3295.3

CG

3614

11.3 90.0

MOMENT

119092.1

CG

MOMENT

-198

-6

485.3 1074.6

0 +559

+9

0 +503

+305

3396.6

3552

120652.0

823.7

3622

29834.4

Takeoff Condition

4220.3

3566

150486.4

+5

+214


3396.6

3552

120652.0

+9

+305

752.8

3559

26792.2

-12

-90


4149.4

3553

147444.2

+5

+215


3396.6

3552

120652.0

+9

+305

345.6

3909

13509.5

0

0

Most Aft Condition

3742.2

3585

134161.5

+8

+305


3396.6

3552

120652.0

+9

+305

147.3

3552

5232.1

0

0

3543.9

3552

125884.1

+9

+305


+Takeoff Fuel (1010.7 Liters)

+Forward Fuel (923.7 Liters)

+Critical Aft Fuel (424.1 Liters)

+Landing Fuel (180.7 Liters)

Landing Condition

-91

-11

BHT-412-FMS-29.2 AND 29.3

MANUFACTURER'SDATA

T Section 2 N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

INSTRUMENT CONSOLES

PANEL

AND

When the upper aft center fuel cell is removed, a 1736 lb fuel capacity placard is mounted on the center section of the instrument panel. See figure 3.

FUELSYSTEM

DESCRIPTION- MECHANICAL

The fuel system (figure 4) is comprised of 9 crash resistant fuel cells. Six of the cells are located below the cabin floor and three are located aft of the cabin and above the level of the underfloor

cells. Refer to figure 5 for fuel burn sequence. Partial cell dividers in two of the aft cells and the system interconnect valve provide 52.5 gallons (198.7 liters) isolated fuel supply for each engine. FUELTRANSFERAND FILLING

Each lower fuel cell is joined with its opposite (left and right), and with the upper cell interconnect system. FUELQUANTITY SYSTEM

The DIGITS TEST button inoperative.

is functionally

13

BHT-412-FMS-29.2 AND 29.3

MANUFACTURER'SDATA


MANUFACTURER'SDATA

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Figure 4.

BHT-412-FMS-29.2 AND 29.3

Fuel system schematic (Sheet 1 of 2)

15

MANUFACTURER'SDATA

BHT-412-FMS-29.2 AND 29.3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O ENGINE 2

Figure 4.

Fuel system schematic (Sheet2)


BHT-412-FMS-29.2 AND 29.3

Figure 5.

Fuel burn sequence

17/18


BHT-412-FMS-29.2 AND 29.3

MANUFACTURER'SDATA

SERVICING

FUELSYSTEMSERVICING

Total capacity: 274.7 U.S gallons (1037.0 liters).

Usablefuel: 267.7 U.S. gallons (1010.7 liters).

19/20

T N I R P N O R I ROTORCRAFT O T F C MANUAL FLIGHT U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAN-30

MODELS412/412EP

CANADIAN ADDENDUM TO THE SUPPLEMENTS FOR

INTERNAL HOIST

AND

EXTERNAL CARGO OPERATION APPROVED FOR DOT NOVEMBER 9, 1989

This addendum shall beattached to the Models 412 and 412 EP Flight Manual when Internal Hoist or External Cargo Suspension Unit is installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this addendum, consult the basic Flight Manual.

Bell Helicopter TEXTRON Inc of Textron A Subsidiary

NOTICE COPYRIGHT

POSTOFFICE BOX 482. FORTWORTH, TEXAS76101

COPYRIGHT 1994

TEXTRON HELICOPTER BELL AND ADIVISION OFTEXTRON CANADA LTD ALLRIGHTS RESERVED

REISSUE -

23 JUNE 1994


BHT-412-FMS-CAN-30

FAA APPROVED SUPPLEMENT FORDOT

LOG OF REVISIONS

Original ..... Reissue.....

0 ..... ......

09 Nov 89 23 Jun 94

LOG OF PAGES Revision No.

Page

Title ............................

0

1/2 .............................

0

A..... .... .... .... ..... .... .. . ..0

NOTICE

This addendum provides additional information to that contained within the Internal Hoist and External Cargo Hook Operation FAA Flight Manual Supplements.

APPROVED:

MANAGER


NOTE: Revisedtext is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

Thispageappliesto CanadianRegisteredhelicoptersonly.

A

T N I SectionI PR N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT FORDOT


Section 1

BHT-412-FMS-CAN-30

FLIGHTCREW LIMITATIONS

HEIGHT-VELOCITYLIMITATIONS

No person shall be carried during internal hoisting or external cargo hook operations unless that person:

The Height-Velocity limitations of the basic manual are not limitations for hoisting or external cargo hook operations.

a. is a crewmember; b.

is a crewmember trainee; or

c.

performs a function operation.

essential to the

Thispageappliesto CanadianRegisteredhelicoptersonly.

T N I R P N ROTORCRAFT O R I O T FLIGHT MANUAL F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-31

BELL BELL MODEL412

SUPPLEMENTFOR CATEGORYB OPERATIONS WITH APPROVEDCONFIGURATION OF NINE OR LESS PASSENGERSEATS

SUPPLEMENTALTYPE CERTIFICATENO. SH7727SW CERTIFIED FEBRUARY8, 1990

This supplement shall be attached to the Model 412 Flight Manual when helicopter is equipped with an approved nine or less passenger seat configuration. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual.

Bell Helicopter

TEXTRON


COPYRIGHT 1990

BELL HELICOPTER TEXTRON INC. ALLRIGHTSRESERVED

FEBRUARY 8, 1990


BHT-412-FMS-31

LOG OF REVISIONS

Original .....

0 ...

.08 FEB90

LOG OF PAGES

Revision No.

Page

Title ........................ A .......................... i/ii .......................... 1-2 ......................... 3/4 .........................

Page

Revision No.

0 0 0 0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATIONADMINISTRATION FT.WORTH, TX 76193-0170

NOTE: Revisedtext is indicated by a black vertical line. Insert latest revision pages. Dispose of superseded pages.

A



BHT-412-FMS-31

INTRODUCTION:

This supplement removes the Height-velocity diagram as a limitation when helicopter is equipped with an approved configuration of nine or less passenger seats.


i/ii

T N Section 1 I R P N O R I O T F C U T O OD N R E P E Section2 E R R C S N G O N I R W O E LY I FSection V 3 N O FAA APPROVED SUPPLEMENT

BHT-412-FMS-31

TYPE OF OPERATION


Flight shall be conducted in accordance with Category B operations and an approved configuration of nine or less passenger seats.

Actual weight change shall be determined after approved seating is installed, and ballast shall be readjusted (if necessary) to return empty weight CG to within allowable limits.

ALTITUDE LIMITATIONS

Maximum altitude for takeoff and landing is 9000 feet density altitude.

TAKEOFFAND LANDING NOTE

The Height-velocity diagram does not represent a limitation. Refer to Performance Data, Section 4.


Useor disclosureof data containedon this pageissubject to the restrictionon the title pageof this document.

1

T N Section 4 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

BHT-412-FMS-31

HEIGHT-VELOCITYENVELOPE

Operation in height-velocity envelope is criticalal in the event of a single engine failure during takeoff, landing, or other operation near the surf ace (figure 4-1). The AVOID area of the the defines diagram Height-velocity combinations of airspeed and height above ground from which a safe single engine landding on a smooth, level, firm surface cannot be assured.

2

The Height-velocity diagram is valid only when the Weight-Altitude-Temperature limitations are not exceeded (refer to basic Flight Manual). The diagram does not define the conditions which assure continued flight following an engine failure nor the conditions from which a safe power-off landing can be made.





FOR SMOOTH, LEVEL, FIRM SURFACES

400 AND ABOVE 375380

120 AND ABOVE 114.3 110

360

340

0

320

300

-90

280

80

260

240

-70

220 200

-60

160

120

60

to

restriction on the title 40 page of this document.

0

10

20

the

30

40

VNE


Figure 4-1.

Height-velocity


diagram (OEI)

3/4

T N BELLMODEL I412 R P N O ROTORCRAFT R I OFLIGHTC T MANUAL F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-33

SUPPLEMENT FOR LORAN C

NAVIGATION SYSTEM (KING KLN-88) 412-899-231 CERTIFIED

22 JUNE 1990


These data are proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation Is forbidden without prior written authorization from Bell Helicopter Textron, Inc. This supplement shall be attached to the Model 412 Flight Manual when the Loran C Navigation System (King KLN-88) has been installed. The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual

Bell Helicopter

POST OFFICE

BOX 482 . FORT WORTH.

BELL'HEUCOPTER TEXTRON INC. AULL RIGHTS RESERVED

TEXAS 76101



LOG OF PAGES

Original .....

0 ....

22 JUNE 1990

FLIGHTMANUAL

MANUFACTURER'SDATA

REVISION NO.

PAGES

Title ......................... A ........................... i/ii .......................... 1 ........................... 2...........................0 3/4 .........................

PAGES

REVISION NO.

0 0 0 0 0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFFICE FEDERALAVIATIONADMINISTRATION FT. WORTH, TX 76193-0170

NOTE: Revisedtest is indicated by a black vertical line. Insert latest revision pages. Dispose of superseded pages.

A



INTRODUCTION

The Loran C Navigation System is a navigation aid for use in the North American geographic area as defined in King KLN-88 Pilot Guide. Visual navigation data, when selected, is presented on the pilot HSI in the form of L/R steering, bearing-to-waypoint, and "TO" indications. L/R steering and "TO" indications are also duplicated on the copilot HSI. The system consists of a combined Loran C receiver and navigational computer, an antenna, a four-way annunciator, switching circuitry and associated wiring.

i/ii

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

TYPE OF OPERATIONS

BHT-412-FMS-33

PLACARDAND DECALS

The Loran C system, as installed in this helicopter, is certified for operation in day or night VFRnon-icing conditions.

OPERATIONALLIMITATIONS

A KLN-88 Pilot Guide (King P/N 006-08458-0000, Operation Revision Status ORS 01) dated August 1989 or later revision. shall be accessible by the flight crew at all times during flight.

The Loran C Navigation shall be operated in accordance with the manufacturers instructions with the following exceptions:

(located on instrument panel)


This Loran C cannot be coupled to the flight director or helipilot.

There is no fuel management data available in this installation.


It is the responsibility of the pilot to verify that any navigation or communications data used is correct.

Actual weight change shall be determined after the Loran C is installed and ballast readjusted as required to return empty weight CG to within allowable limits.

1

T N Section 2 I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-33

EXTERIORCHECK 7.

CABIN TOP


Pilot HSI CRS pointer - Align to desired course shown on Loran display. NAV/LRN switch annunciator -

Press; verify

Loran C antenna - Condition and security.

LRN segment illuminated and NAV segment extinguished.

PRESTARTCHECK

Pilot and copilot HSI deviation bars - Verify centered and "TO" indication displayed.

LORAN PWR and FAN circuit breakers - In. Loran C unit - Verify off. BEFORE TAKEOFF

Loran C unit - Turn on. Verify operational revision status on initial display page is identical to that of available KLN-88 Pilot Guide.

2

Pilot HSI bearing pointer - Verify bearing to waypoint displayed. NOTE

For additional normal procedures, except fuel management, refer to KLN-88 Pilot Guide.

T N I R P N O R I O T Section F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-33


NO CHANGE

NO CHANGE

3/4

T N I BELL MODEL R 412 P N ROTORCRAFT * O R I O T FLIGHTMANUAL F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-34.2

33108 THROUGH 33213 AND 36001 THROUGH 36019

SUPPLEMENTFOR IMPROVEDHOVER PERFORMANCE WITH PT6T-3BEENGINES AND 5-MINUTE TAKEOFFPOWERRATING (412-570-001-103) CERTIFIED OCTOBER12, 1990

PROPRIETARYRIGHTSNOTICE

These data are proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron, Inc. This supplement shall be attached to the Model 412 Flight Manual when the Improved Hover Performance Modification (412-570-001-103) has been installed.

The information contained herein supplements the information of the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual



LOGOF PAGES

Original .... Reissued ..

0 ....... ...........

12 October 1990 8 April 1991

MANUFACTURER'SDATA

FLIGHT MANUAL

REVISION NO.

PAGES

Title ......................... A ........................... ................... 1 -40 41/42 ....................... 43 - 44 ...................

PAGES

0 0 .0 0 .0

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

NOTE: Revised text is indicated by a black vertical line. Insert latest revision pages. Dispose of supersededpages.

REVISION NO.

T * Section 1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

BHT-412-FMS-34.2


AIRSPEEDLIMITATIONS

Actual weight change shall be determined after components are installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits.

Airspeed shall not exceed 105 KIAS (or placarded VNE, if less) when operating above 81% mast torque.

Weight-Altitude-Temperature to Refer Limitations chart for maximum allowable weight for takeoff, landing, and IGE hover operation.

Refer to Maximum Speed-Sideward and Rearward Flight, Crosswind and Tailwind At A Hover chart.

1



NOTE: ALLOWABLEGROSSWEIGHTS OBTAINEDFROMTHIS CHART MAY EXCEED CONTINUOUSHOVERCAPABILITYUNDERCERTAINAMBIENT CONDITIONS. REFERTO HOVER CEILINGCHARTS IN SECTION4.

9

4.0

10

11

12 LBX1000



NOTES: 1. Pedal critical wind azimuth-hovering with the relative wind within these azimuth angles

0

hardover

IGE CRITICALRELATIVEWIND AZIMUTH


14,000 FT DENSITY ALTITUDE

12,000

.


BHT-412-FMS-34.2

POWERPLANT LIMITATIONS

Pratt and Whitney Aircraft of Canada, Ltd. PT6T-3BE. POWER TURBINE RPM (Nm LIMITS

97%

Minimum in cruise

ATTITUDE CHANGES. DECREASE POWER TO 90% MAST TORQUE MANEUVERING TO PRIOR HELICOPTER.


Minimum for hover, takeoff, and climb

100%

TRANSMISSION TORQUELIMITS

Maximum continuous

100%

Deleted. See MAST TORQUELIMITS.

MAST TORQUELIMITS

ROTORLIMITATIONS

TWIN ENGINE OPERATION Maximum continuous

Takeoff range (5 minutes maximum)

ROTOR RPM (Nr) LIMITS - POWERON

81%

81 to 100%

97%

97 to 100%


Maximum continuous

WARNING

TAKEOFF POWER SHALL NOT BE USEDABOVE 105 KIAS.

100%

Maximum

Minimum

100%

Operation with mast torque at or below 32%

100 to 104.5%

Maximum with mast torque at or below 32%

104.5%

CAUTION

WHEN OPERATING NEAR THE MAXIMUM MAST TORQUE LIMIT, INADVERTANT OVERTORQUE MAY OCCUR DURING MANEUVERING FLIGHT CONDITIONS INVOLVING TURNS AND/OR NOSE DOWN

FUEL AND OIL LIMITATIONS

TRANSMISSION, INTERMEDIATE AND TAIL ROTORGEARBOXOIL

Turbine Oil 555 is the only approved oil for use in the transmission and gearboxes.

TRIPLETORQUEINDICATOR MAST TORQUE

GREEN YELLOW RED

10to81% 81 to 100% 100%

Continuous operation 5 minute takeoff range Maximum

ENGINE

GREEN YELLOW RED

5 to 58.9% 58.9 to 73.2% 73.2%

Instrumentmarkings

Continuous OEI operation 30 minute OEI range Maximum OEI

T * Section 2 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N Section 3 I R W O E I F V NLY O BHT-412-FMS-34.2


EXTERIORCHECK

RPM switch seconds).

FUSELAGE - AFT LEFT SIDE

RPM switch seconds).

Check OVER TRQ warning flag (cat's eye) for indication of overtorque.

Minimum beep (DECR for 4-5

-

Minimum trim (-2 for 4-5

NR - Check 95% or greater.

PRESTARTCHECK

RPM switch - Adjust to obtain matching torque at 100% NR.

OVERTORQ caution light - Press. Check light illuminates and MAST TORQUE indicator reads 105±1%.

Flight instruments - Check operation and set.

CAUTION

IF MAST TORQUE INDICATOR INDICATES AN ERROR GREATER THAN ± 1% FROM THE 105% POSITION, THE MAST TORQUE SYSTEM IS UNRELIABLE. MAINTENANCE ACTION IS REQUIRED.

BEFORETAKEOFF

Throttles - Full open. Adjust frictions.

TAKEOFF

Area - Clear.

NOTE As collective is increased, it may be necessary to rematch engine torques prior to reaching hover.

RPM switch - Adjust to obtain matching torque or ITT, as required, and 100% NR.

Hover power - Check torque required to hover at four feet skid height.

EMERGENCYPROCEDURES Caution lights

PANEL WORDING OVER TORQ

FAULTCONDITION

CORRECTIVEACTION

Mast torque exceeds 100.5%

Reduce power or severity of maneuver. Land as soon as practical.

T N Section 4 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED SUPPLEMENT

BHT-412-FMS-34.2

I

INTRODUCTION

The performance data presented herein are derived from the engine manufacturer's specification power for the engine less installation losses when used with the 412-570-001-103 Improved Hover Performance modification. These data are applicable to the basic helicopter without any optional equipment which would appreciably affect lift, drag, or power available.

HOVERCEILINGOGE.

AREA A (unshaded area) as shown on the hover ceiling charts presents hover performance for which satisfactory cyclic and directional control

have been demonstrated in relative winds of 35 knots from any direction at or below 3000 feet HD. Improved control margins will be realized by avoiding winds in the critical relative wind azimuth areas.

AREAB (shaded area) as shown on hover ceiling charts presents additional hover performance which can be realized in calm winds or winds outside the critical relative wind azimuth areas. NOTE

Tail rotor or cyclic control margin may preclude operation in AREA B of the hover ceiling charts when the relative wind is in the respective critical wind azimuth area.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SEENOTE

NOTES: 1. Pedalcritical wind azimuth-hovering with the relative wind within these azimuth angles

7


Hover ceiling (Sheet 1 of 10)

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER CEILING


ENGINE TAKEOFF POWER ENGINE RPM 100% GENERATOR 150 AMPS IEA.)

CAUTION:

OGE HOVER OPERATION MAY RESULT IN VIOLATION OF H-V LIMITATIONS.


9



HOVERCEILING


ENGINE TAKEOFF POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

CAUTION:

SKID HEIGHT 60 FT. HEATER OFF -40 TO 0C


--

14.000 FT.DEN. ALT.LIMIT

OAT- C

8

9


10

11

12

LBx1000


BHT-412-FMS-34.2

HOVER CEILING



ENGINE TAKEOFF POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

CAUTION:


FT. DEN. ALT. LIMIT

14,000

AREA A

0 10 20 OAT- C

8

9

10

11

12

LBx1000


11



HOVERCEILING

OUTOF GROUNDEFFECT

ENGINETAKEOFFPOWER ENGINERPM 100% GENERATOR 150 AMPS(EA.)

SKID HEIGHT 60 FT. HEATER ON -40 TO 0°C

MAY RESULT IN VIOLATIONOF H-V LIMITATIONS. CAUTION: OGEHOVEROPERATION

FT.DEN. ALT.LIMIT

AREA A



BHT-412-FMS-34.2

HOVER CEILING IN GROUND EFFECT

MAXIMUM CONTINUOUS POWER ENGINERPM 100% GENERATOR150 AMPS (EA.)

SKID HEIGHT 4 FEET HEATER ON OR OFF -40

TO 52°C

MAX

MAX OAT HEATER ON (21°C)

-40-30-20-10

0 10 20 OAT - °C

30

40

50 60

8

3.5

9

10

11


12 LB X 1000

5.4 kg X 1000


13


BHT-412-FMS-34.2

HOVER CEILING OUT OF GROUND EFFECT

SKID HEIGHT 60 FT. HEATEROFF 0 TO 52°C


MAY RESULTINVIOLATIONOF H-VLIMITATIONS. CAUTION: OGEHOVEROPERATION 14,000 FT. DEN. ALT. LIMIT

9

Hoverceiling (Sheet 7 of 10)

10

11

12 LB X 1000


BHT-412-FMS-34.2


MAXIMUM CONTINUOUS POWER ENGINERPM 100% GENERATOR150 AMPS (EA.)

SKID HEIGHT 60 FT. HEATER OFF -40 TO 0°C

CAUTION: OGEHOVEROPERATIONMAY RESULTINVIOLATIONOF H-VLIMITATIONS.

14,000 FT.DEN. ALT. LIMIT AREA B

OAT -C

Hoverceiling (Sheet8 of 10)

15


BHT-412-FMS-34.2



SKID HEIGHT 60 FT. HEATERON 0 TO 20°C

MAY RESULTINVIOLATIONOF H-VLIMITATIONS. CAUTION: OGEHOVEROPERATION





MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.) CAUTION:



-

14,000 FT DEN. ALT. LIMIT-\

•-,

)-30-20-10

0

8

9

-- TORQUE LIMIT

10

11

12 LB X 1000


17


ENGINERPM 100% GENERATOR 150 AMPS(EA.)

WITH ALL DOORSOPENORREMOVED: 1. CLIMBSPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT./MIN. 20,000

18.000

16.000

14.000


70 KIAS HEATERON


TAKEOFFPOWER ENGINERPM 100% GENERATOR 150 AMPS(EA.)

TWIN ENGINE RATE OF CLIMB GROSSWEIGHT11,900 LB (5398 kg)

WITH ALL DOORSOPENOR REMOVED: 1. CLIMBSPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT./MIN.

18,000

16,000

-4500

14.000

-4000

\


19



TWIN ENGINE RATEOF CLIMB

TAKEOFFPOWER ENGINERPM 100% GENERATOR 150 AMPS(EA.)

GROSSWEIGHT11.000 LB(4990 kg)

70 KIAS HEATERON

WITH ALL DOORSOPENORREMOVED: 1. CLIMBSPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT./MIN.

-6000

-5500

-5000

-4500

0

400


BHT-412-FMS-34.2

TWIN ENGINERATEOF CLIMB

GROSS WEIGHT 11,000 LB (4990 kg)



70 KIAS HEATEROFF

1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT./MIN.

20,000

- 6000

- 5500

18,000

- 5000

-4500

-4000


21


TWIN ENGINERATEOF CLIMB

GROSS WEIGHT 10.000 LB (4536 kg)

70 KIAS HEATER ON




20,000

-6000

18,000

-5500

-5000

16.000

-4500

14,000

- 4000


22


BHT-412-FMS-34.2

TWIN ENGINE RATE OF CLIMB GROSSWEIGHT10,000 LB (4536 kg)

TAKEOFFPOWER ENGINERPM 100% GENERATOR 150 AMPSIEA.)

70 KIAS HEATEROFF

WITH ALL DOORSOPENORREMOVED: 1. CLIMB SPEEDIS 60 KIAS 275 FT./MIN. 2. RATEOF CLIMB WILLDECREASE 20,000

-6000

0

400

(0)

(2.0)

800

1200 1600 2000 2400 RATE OF CLIMB - FEET/MINUTE

(4.0) (6.0) (8.0) (10.0) (12.0) RATE OF CLIMB - (METERS/SECOND)

2800

(14.0)

3200

(16.0)


23



TWIN ENGINE RATEOF CLIMB GROSSWEIGHT9000 LB (4082 kg)

TAKEOFF POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.I

70 KIAS HEATER ON



20,000

- 6000

OAT

-5500

- 5000

-4500

- 4000

-3500

-3000

-2500

-2000

-1500

-1000

- 500

0

0

400

800

(0

(2.0)

(4.0)

1200 1600 2000 2400 RATE OF CLIMB - FEET/MINUTE (6.0)

(8.0)

(10.0)

(12.0)

RATE OF CLIMB - (METERS/SECOND) Twin engine rate of climb (Sheet 7 of 24)

2800

(14.0)

3200

(16.0)


BHT-412-FMS-34.2

TWIN ENGINERATEOF CLIMB GROSSWEIGHT9000 LB (4082 kg)


70 KIAS

HEATEROFF


1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE275 FT./MIN.

20,000

-6000

18.000

-5500

- 5000

16,000

-4500

14,000

- 4000

12.000

-3500

u-

-3000

OAT LIMIT

-2500

-2000

-1500

-1000

RATE OF CLIMB - (METERS/SECOND) Twin engine rate of climb (Sheet 8 of 24)

25



70 KIAS HEATER ON

20.000

18,000

16,000


26



TWIN ENGINE RATE OF CLIMB GROSSWEIGHT8000 LB (3629kg)




18,000

16,000

14,000

u

- 500


27



TWIN ENGINE RATE OF CLIMB GROSSWEIGHT7000 LB (3175 kg)

TAKEOFFPOWER ENGINERPM 100% GENERATOR 150 AMPS (EA.)

70 KIAS HEATERON

WITHALL DOORSOPENOR REMOVED: 1. CLIMBSPEEDIS 60 KIAS 2. RATEOF CLIMBWILLDECREASE 275 FT./MIN.

20.000

-6000

OATLIMIT

-5500

18.000

- 5000

16,000

- 4500

14,000

-4000

w

-3500

10,000

-3000 -2500

8,000

- 2000

6,000

-1500

4,000

-1000

2.000

- 500

0

0

400

800

(0)

(2.0)

(4.0)

1200 1600 2000 2400 2800 RATEOF CLIMB - FEET/MINUTE

(6.0) (8.0) (10.0) (12.0) (14.0) RATEOF CLIMB - (METERS/SECOND)


3200

(16.0)

3600

(18.0)


BHT-412-FMS-34.2

TWIN ENGINERATEOF CLIMB GROSS WEIGHT 7000 LB (3175 kg)



70 KIAS HEATER OFF


-6000

-5500

-5000

-4500

1000



BHT-412-FMS-34.2

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT

7000

LB (3175

kg)

70 KIAS HEATER OFF


WITH ALL DOORS OPEN OR REMOVED

1 CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT /MIN.

20,000

18,000

16,000

Twinenginerate of climb (Sheet13of 24)


BHT-412-FMS-34.2


MAXIMUM CONTINUOUS POWER ENGINERPM 100% GENERATOR 150 AMPS (EA.)

70 KIAS HEATER ON

WITHALL DOORSOPENOR REMOVED 1 CLIMBSPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT/MIN


31


BHT-412-FMS-34.2

TWIN ENGINE RATE OF CLIMB

GROSS WEIGHT 8000 LB (3629 kg)

70 KIAS HEATER OFF

MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR150 AMPS (EA.)

WITHALL DOORSOPENOR REMOVED 1 CLIMBSPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT /MIN

20,000

\

6000

RATE OF CLIMB

RATE OF CLIMB -

-

FEET/MINUTE

METERS/SECOND)


BHT-412-FMS-34.2


GROSS WEIGHT 8000 MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR 150 AMPS (EA.) WITH ALL DOORS OPEN OR REMOVED

LB (3629 kg)

70 KIAS HEATER ON

1 CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT /MIN.

20,000

16.000

14,000


33




GROSS WEIGHT 9000 LB (4082 kg) 70 KIAS MAXIMUM CONTINUOUS POWER HEATEROFF ENGINE RPM 100% GENERATOR150 AMPS (EA.) WITHALL DOORSOPENOR REMOVED. 1 CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT./MIN

20.000

18,000

16,000

Twinenginerateof climb (Sheet17 of 24)

34


20,000

18,000

16.000

I-

Lu Lu

6.000

BHT-412-FMS-34.2




GROSS WEIGHT 10.000 LB (4536 kg) MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR150 AMPS (EA.)

70 KIAS HEATEROFF

WITHALL DOORSOPENOR REMOVED 1. CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT /MIN.

20,000 18.000

-I

5500

-!

5000

-

1500

16.000 "

I-

uJ

_-

-I 6000

\'

14,000

LU

I oo000ou

12.000

3500 UJ

0

10.000

3000 5 4

LU

,-

uJ

2500 §

8.000

UJ

0.

2000

6.0()0

1500

4,0(

OAT LIMIT-

1000

2.0()O

500

20

o 0

200

40

60

80

10

1

1600

.0 (60 7.0) (.0 (2.0) (3.0) (4.0) (5.0) RATE OF CLIMB - (METERS/SECOND)

(8.0)

600

800

1000

1200

1

1400

400

RATE OF CLIMB - FEET/MINUTE

)

(1.


36

X


BHT-412-FMS-34.2

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB 14536 kg) MAXIMUM CONTINUOUS POWER 70 KIAS ENGINE RPM 100% HEATER ON GENERATOR150 AMPS (EA.) WITHALL DOORSOPENOR REMOVED 1 CLIMBSPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT/MIN

20,000

-6000 -5500

-20C

16 ,000

-5000

40C

-4500

14,000

- 4000

12,000

-3500

10,000C

- 3000

8,000

-2500

- 2000

-1500

4,000

-1000

2,000

-500

0

0

200

(0)

(1.0)

400 600 800 1000 1200 RATE OF CLIMB - FEET/MINUTE


1400

(7.0)

1600

(8.0)


37


BHT-412-FMS-34.2



MAXIMUM CONTINUOUSPOWER 70 KIAS ENGINE RPM 100% HEATER OFF GENERATOR 150 AMPS (EA.) WITHALL DOORSOPENOR REMOVED. 1. CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMB WILL DECREASE 275 FT./MIN.

20.000

-6000 -5500

18.000

OAT OAT

16,000

-5000 -4500

-20C

14,000

UJ

-4000

12.000

I

- 3500

0 I

4

10,000

- 3000

-2500

8,000

- 2000

6,000

-1500

4.000

I

OAT

LIMIT

1000

2,0(

\

0

0

200

400

600

800

1000

500

1200

1400

1600


(0)

(110)

(2.0)

(3.0)

(4.0)

(5.0)

(6.0)

RATE OF CLIMB - (METERS/SECOND) Twinenginerate of climb (Sheet21 of 24)

38

(7.0)

(8.0)


BHT-412-FMS-34.2

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 11,000 LB (4990 kg) MAXIMUM CONTINUOUS POWER 70 KIAS ENGINE RPM 100% HEATERON GENERATOR 150 AMPS (EA.) WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMB WILL DECREASE 275 FT /MIN

20.000

-6000

18.000

-5500

OAT

6,000

16,000

-5000

30C

-4500

I-

14,000

-4000

12.000

w

-3500

10.000

- 3000

8,000

-2500

I

,

-2000

6,000

-1500

4,000

-1000

2.000

-500

0 0

200

(0)

(1.0)



1400

(7.0)

1600

(8.0)


39


BHT-412-FMS-34.2


GROSS WEIGHT 11.900

LB (5398 kg)

70 KIAS HEATER OFF



1 CLIMB SPEED IS 60 KIAS 2 RATE OF CLIMB WILL DECREASE 275 FT./MIN.

20,00

6000

OAT

5500

500

-5 5000

40°C -20C

4500

I-

UJ

4000 3500 1500

1

2500

1500

OAT

-11000

LIMIT

500

0

200

(0)

(1.0)


1400

(7.0) (6.0) (5.0) (4.0) (3.0) (2.0) RATE OF CLIMB - (METERS/SECOND) Twinenginerateof climb (Sheet23 of 24)

1600

(8.0)


BHT-412-FMS-34.2


GROSS WEIGHT 11,900 LB (5398 kg) MAXIMUM CONTINUOUS POWER ENGINERPM 100% GENERATOR 150 AMPS (EA.)

70 KIAS HEATERON

WITHALL DOORSOPENOR REMOVED 1 CLIMBSPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT /MIN

-6000 -5500 -5000 -4500

14,000

LL

-4000

12,000:

I

-3500

10.000

- 3000

-2500

8,000

0.

-2000

6,000

-1500

4,000 4,000

1000

2.000

500

0

0

200

(0)

(1.0)


1400

(210) (3.0) (4.0) (5.0) (6.0) (7.0) RATE OF CLIMB - (METERS/SECOND)

1600

(8.0)


41/42

T N I Section 1 R P N O R I O T F C Section 2 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-34.2

MANUFACTURER'SDATA


MANUFACTURER'SDATA

INSTRUMENT CONSOLES

PANEL

AND

AMBER


OVERTORQUEWARNING SYSTEM

An OVER TRQwarning flag (cat's eye) is located in the lower left aft avionics compartment. If mast torque exceeds 108%, the warning flag will be tripped showing alternating black and white sections.

POWERPLANT

The pilot RPM switch is mounted on the collective switchbox. The pilot switch is a five-position momentary-on type switch. The INCR position increases engine RPM. and the DECR position decreases engine RPM. The INCR/DECRpositions control the governors on both engines simultaneously. Regulated engine RPM may be adjusted in flight through the operating range of 97 to 101.5 (±0.5) % by moving the switch.

(ReplacesTRANSMISSIONTORQUEgage)

The RPM +2/-2 switch increases or decreases engine No. 2 RPM to provide torque or ITT matching. Engine2 trim range is 2.0 - 2.5% NII,. Engine 2 governor should be at least 95% minimum trim - minimum beep.

43

T N I R P N O R I O T F C U T O OD N R E P E E R R Section 3 C S N G O N I R W O FSection 4 VIE LY N O BHT-412-FMS-34.2

The copilot does not have trim capability. For location of RPM switch, refer to collective control panel.

ON

ARMED

EXT

G L

OFF

START

IDLE STOP

ENG 2

1

INCR

ENG

SRCH LT

Collective control panel.

MANUFACTURER'SDATA


MANUFACTURER'SDATA


F T

T N mell F I R P N ROTORCRAFT O I 0) OR FLIGHT MANUAL T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-35.3 AND 35.4

MODELS

SUPPLEMENT

CATEGORY B OPERATIONS WHEN CONFIGURED WITH NINE OR LESS PASSENGER SEATS 36020 -

36086

36087 AND SUBSEQUENT AND 33108-33213 36001 -36019 WHEN 412-570-001-103 MODIFICATION (BHT-412-FMS-34.2) IS INSTALLED CERTIFIED 10 APRIL 1991

This supplement shall be attached to Model 412/412EP Flight Manual when helicopter is configured with nine

or less passengers seat configuration.

Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, or other applicable supplements, consult basic Flight Manual.

COPYRIGHT NOTICE COPYRIGHT 2000 ® BELL

HELICOPTERTEXTRONINC.

Bell Helicopter *Lit ASubsidiary of Textron Inc POST OFFICE BOX 482

FORT WORTH. TEXAS 76101

ANDBELLHELICOPTER TEXTRON INC. A DIVISION OF TEXTRONCANADALTD.

ALLRIGHTS RESERVED5

REVISION 3 -5

10 MAY R1996

SEPT

SEPTEMBER 2000


FAA APPROVED

NOTICE PAGE

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 76101-0482

NP

SEP 052OO

BHT-412-FMS-35.3 AND 35.4

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODELS

/b

ROTOR CRAFT FLIGHT MANUAL SUPPLEMENT

RY B OPERATIONS WHEN JED WITH NINE OR LESS

SENGER SEATS

36020

D 36087 AND SUBSEQUENT D-.86 /j// /%WAND 33 r /'.ZRAND -36001 - 36019

WHEN TH/4

MODIFICATION

1-103

(BHT4

IS INSTALLED E4

TEMP RYyISION

MAST TORQUE SIly4ONDITIONER

ASB

This Tmporay

Revsion

exch

cunpubished t

newt perm anm d pi comnpliace with ASB 4127-0

aler

\

Insert these temporary revision pages pages in Right Manual after compliance

u

DO NOT remove existing pages from Flight M

infortnaon ainod rein supplemns RiFt Manu. For miations, Procedures, and P Data not contained in this supplement consult basic Right

Manual.


Bell HellcopteritwiI

EBXir_

HTIA CO ym 4 TEXIMM 'fA 4C.

* L T''"OmoA mL. urnnmrlMGHS ESE

REISSUE -10

tiy

11

01Tu-on

Ft

MAY 1996


I


LOG OF REVISIONS

Original........... 0 .............. 10 APR 91 Reissue...........0 ............. 10 MAY 96 Revision .......... 1 ............. 08 AUG 97

*^&g~~

~LOG

OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title .................................... NP ...................................... A- B .................................. i/ii .......................................

*^&-~~~

Revision .......... 2 .............. 23 APR 98 Revision .......... 3 .............. 05 SEP 00

3

3 0

REVISION NO.

PAGE

1 ........................................ 2 ........................................ 3-5 ................................... 6- 9 ................................... 10 ....................................... 11-12.................................0

0 3 0 3 2

~~NOTE


SEP 05 2000

Rev. 3

A

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-35.3 AND 35.4


Temporary ...................... 8 AUG 97 Canceled ....................... 5 SEP 00

NOTICE

This Temporary Revision has been cancelled. Subject data has been incorporated in BHT-412-FMS-35.3 and 35.4 at Revision 3. Remove and discard all Temporary pages dated 8 August 1997 from this manual including these canceled pages.

APPROVED:

SEP

5 200

MANAGER

ROTORC AFT CERTIFICATION OFFICE FEDERA AVIATION ADMINISTRATION FT. WORTH, TX 761934170 B

TEMPORARY REVISION - 8 AUG 97 CANCELED - 5 SEP 2000



Original...........0 .............. 10 APR 91 Reissue...........0 ............. 10 MAY 96 Revision..........1 ............. 08 AUG 97

Revision .......... 2 .............. Revision .......... 3 ..............

DATE:

APPROVED:

SEP

23 APR 98 05 SEP 00

5 2000

MANAGER

ROTORCRT CERTIFICATION OFFICE FEDERA LAVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

B

Rev. 3

SEP 0 5 200

0


BHT-412-FMS-35.3 AND 35.4

GENERAL INFORMATION

This supplement removes the Height-velocity diagram as a limitation and increases the density altitude limit for takeoff, landing, and in ground effect maneuvers when helicopter I is configured with nine or less passenger seats.

i/ii

T N I * Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-35.3 AND 35.4

FAA APPROVED

4

1-3.

TYPES OF OPERATION

Flights may be conducted in accordance with this supplement only when the helicopter is configured with nine or less passenger seats.

1-6. WEIGHT AND CENTER OF GRAVITV--------Y nGAVIITYXlanding,

Maximum gross weight for takeoff, landing, and in-ground-effect maneuvers is 11,900 pounds (5,398 kilograms) or as shown in Hover ceiling in ground effect

(takeoff power) chart, refer to Section 4, whichever is less.

Actual weight change shall be determined after seating is installed, and ballast shall be adjusted (if necessary) to return empty weight CG to within allowable limits.

1-8. ALTITUDE

Maximum density altitude for takeoff,

and in-ground-effect maneuvers is

16,000 feet. Refer to Weight-AltitudeTemperature Limitations chart (figure 1-1).


FAA APPROVED

NOTE: ALLOWABLE GROSS WEIGHTS OBTAINED FROM THIS CHART MAY EXCEED HOVER CAPABILITY UNDER CERTAIN AMBIENT CONDITIONS. REFER TO HOVER CEILING CHARTS IN SECTION 4.


2

___^

--

SEP 05 2000

x^^^

Rev. 3

--

//

/;^^

/zz -Irr _ - _ _ _zz~ ~/LGz_ _ -^ i/y-^ -~~~v.~,,"/'- -_ _ -_ -_ I

-40

I

I

-20

0 OAT

--

~-^

20

I y

40

60

, /GROSS

-

--

-

8 9 10 12 GROSS WEIGHT LBS X11x1000 1000 WEIGHT -- KG

412FS-35.3-1 -1

Figure 1-1.

Weight-altitude-temperature-limitations for takeoff, landing and in-groundeffect maneuvers.

2 Rev. 3 SEP 052000

T N * Section 2 I R P N O R I O T F C U T D O Section O 3 N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-35.3 AND 35.4

FAA APPROVED

*


Refer to Performance Data, Section 4.

NOTE

The Height-velocity diagram does not represent a limitation.


3

T N I Section 4 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-35.3 AND 35.4

4-4. HEIGHT ENVELOPE

-

VELOCITY

Operation in height-velocity envelope is critical in the event of a single engine failure during takeoff, landing, or other operation near the surface (figure 4-1). The AVOID area of the Height-velocity diagram defines the combinations of airspeed and height above ground from which a safe single engine landing on a smooth, level, firm surface cannot be assured. The Height-velocity diagram is valid only when the Hover Ceiling Out of Ground Effect performance envelope is not exceeded (refer to figure 4-2). The Hover Ceiling In Ground Effect performance chart (refer to figure 4-3) does not define the conditions which assure continued flight following an engine failure nor the conditions from which a safe power-off landing can be made.

4-5. HOVER CEILING

The Hover Ceiling In Ground Effect charts (figures 4-3 and 4-4) provide the maximum allowable gross weights for takeoff,

4

FAA APPROVED

landing, and IGE maneuvers at all pressure

altitude and outside air temperature conditions with heater on or off. Conversely, the hover ceiling altitude can be determined for any given gross weight.

Adequate cyclic and directional control are available at the gross weights allowed by the Hover Ceiling IGE charts in winds up to 35 knots from any direction at or below 3000 feet HD (refer to Basic Flight Manual). Above 3000 feet HD, improved control margins will be realized by avoiding winds in the critical wind azimuth area (figure 45).

4-6.

TAKEOFF DISTANCE

The Takeoff Distance charts (figure 4-6) provide takeoff distances required to clear a 50 foot or 15 meter obstacle in a zero wind condition. Takeoff is initiated from a hover at 4 feet (1.2 meters) skid height with climbout of 45 knots. NOTE

Downwind takeoffs are not recommended because the published takeoff distance performance cannot be achieved.


BHT-412-FMS-35.3 AND 35.4


FOR SMOOTH. LEVEL. FIRM SURFACES

400 ABOVE AND AND AOVE 37

I

NOTE:

--

120

I12

THE HEIGHT-VELOCITY DIAGRAM IS VALID ONLY WHEN THE HOVER CEILING OUT OF GROUND EFFECT PERFORMANCE . ENVELOPE IS NOT EXCEEDED (REFER TO -4 FIGURE 4-2).

380 360

AND ABOVE 114.3 110

340

*"

320 3_0--100

------

1.1 e

300

uL

280

Q

2602

S

g

_

2

_ 4o _

_

_

240

I

220

iI---

0

200 o

i0-

<

2

,

^

^

1a1200

10

160

___

^

20

L-----_-_-_

30-40

0

fI

-50

9o I

4

INDICATED AIRSPEED100 - KNOTS 8i t80S8L g 40 ei ecity dia

^^&

-70

-30

-70 o5,-I-_--------------. 9

10

20

30

40

INDICATED AIRSPEED-KNOTS

VME

-

T N I R P N O 0 R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-35.3 AND 35.4

FAA APPROVED



SKID HEIGHT 60 FEET HEATER ON OR OFF -40 TO 0° C

NOTE: THESE DATA VALID FOR ZERO WIND OUTSIDE OF THE CRITICAL WIND AZIMUTH AREA (REFER TO FIGURE 4-5)


MA IMUM

If _AX

-40

-30

-20

-10 -0 0

OAT - QC

8

9

11

12

GROSS WEIGHT - LBS x 1000 I I I I I I I

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

412FS-3.34-2-10 Figure 4-2.

6

Rev. 3

Hover ceiling out of ground effect (Sheet 1 of 2)

SEP 052000

0


BHT-412-FMS-35.3 AND 35.4



SKID HEIGHT 60 FEET HEATER ON OR OFF 0 TO 52°C

NOTE: THESE DATA VALID FOR ZERO WIND OUTSIDE OF THE CRITICAL WIND AZIMUTH AREA (REFER TO FIGURE 4-5)


t

o,/

2

I

^_/_-~/ 3'

I

7T 7_C

7-

IM

_

d /~4-\--------_' I

OAT

/

o

-

7,/

_S

_EIGHT

sEIC

HT^

\\

_--------

/

-S

ES

GROSS WEIGHT -

LBS x 1000


412FS-35.3-42-2

Figure 4-2.

Hover ceiling out of ground effect (Sheet 2)

SEP 05 2000

Rev.3

7


FAA APPROVED

HOVER CEILING IN GROUND EFFECT MAXIMUM WEIGHT FOR TAKEOFF, LANDING, AND IN-GROUND EFFECT MANEUVERS

TAKEOFF POWER SKID HEIGHT 4 FEET ENGINE RPM 100% HEATER ON OR OFF GENERATOR 150 AMPS(EA.) -40 TO 52°C NOTE: THESE DATA VALID FOR ZERO WIND OUTSIDE OF THE CRITICAL WIND AZIMUTH AREA (REFER TO FIGURE 4-5)


/ '

8/ Rev. 3

/1

_______

SEP / /5 200

M-

WIH

KG

Fiu/re--3 H ve ci ligiron fe ct- -- -- 100 -- -» SEP 05 2000_ __ __ __ ____ __ __ __ ___ Ii/ / .,^ _^ __________ 8 Rev.3

II Y,~,/!-^_/-^~,_ /l/_ _^ /

-40

-20

Figure 4-3. 0 20

___

_

;

_z:: ___

Hover 40 ceiling 60 in ground 8 9effect 10

11

12


8

Rev. 3 $EP 052O00

^


BHT-412-FMS-35.3 AND 35.4

HOVER CEILING IN GROUND EFFECT MAXIMUM WEIGHT FOR TAKEOFF, LANDING, AND IN-GROUND EFFECT MANEUVERS

MAXIMUM CONTINUOUS POWER SKID HEIGHT 4 FEET ENGINE RPM 100% HEATER ON OR OFF GENERATOR 150 AMPS(EA.) -40 TO 52 0 C NOTE: THESE DATA VALID FOR ZERO WIND OUTSIDE OF THE CRITICAL WIND AZIMUTH AREA (REFER TO FIGURE 4-5)


-, /

/~

^

/~t

<

_

A-

-

/

\\X

IOA

T

_S_

I

-CGROSS

AG,

_

IGT

G

\-VI

-

WEIGHT

W/1

-

x

1000

412FS35.3-4-4

Figure 4-4.

Hover ceiling in ground effect

SEP 052000 SEP 05200

Rev. 3 9 Rev.3

9

T N I R P N O R !H I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-35.3 AND 35.4

CRITICAL RELATIVE WIND AZIMUTH

00

330 3153

3

'"

45.

Critical wind azimuth - hovering with the relative wind azimuth can ca angles in shaded area result in the following:

4355

Inability to maintain headingdue SS Ha. to large left pedal requirements for certain wind velocities.

270

b. Reduction of available left pedal control with a directional AFCS hardover. (Excluding 412EP).

c. Aft cyclic may be limited with longitudinal AFCS hardover.

615,000

180' 16,000 FT. H IGE MANEUVERS LIMIT FOR IGE LIMIT

__7

______

14,000

A /2\

12.000 _ I.-

_

_

_

_-

Wu

i

_

_

<7 78

7

7

A_ For HD below 14,000 ft winds up to ±450 off nose for hover performance to be valid.

'//N ,/W

_0

E 8000__ 4.000

/

/

0,.000… -

For H. from 14,000 to 16,000

° winds up to +30 off nose for hover performance to be valid.

78f7/

_---77ft /7

_- -

.

Vi

A

l

d. Hover performance is valid for all headings in calm wind.

-- __E

--

oo

S4

43.000

_NOTE: Refer to Basic Flight Manual for critical wind azimuths at density altitudes at or below

3,000 feet

o0I 14

I

16 18 20 22

I I I II

I I \YI x

24 26 28 30 32 34

36 38

MAXIMUM ALLOWABLE WIND SPEED-KNOTS

Figure 4-5.

10

Rev. 2

Critical relative wind azimuth

412FS35.3 -4-5

0


BHT-412-FMS-35.3 AND 35.4

TAKEOFF DISTANCE OVER 50 FOOT OBSTACLE

HOVER POWER +15% TORQUE (not to exceed limits) ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

---

^9

/16.000

\\/i

FT DEN. ALT. LIMIT.

_ Ii~ 1

1A/

OAT-oC _ A

-40 -o

INITIATED FROM 4FT. SKID HEIGHT VTO = 45 KIAS HEATER ON OR OFF

o

20

OAT--C _

I

-

-

_

40

I

_ANCE

4 40

S

~IIII//IENOLISH

Figure 4-6.

/

GROSSWEIGHT

-

O

FT.

it;

TAKEOFFDISTANCE FT.

Takeoff distance (English)

11

T N I R P N O R I O T F C U T O OD N R E P E E R R OATI C S I; N G O N I R W O E I F I, V NLY O BHT-412-FMS-35.3 AND 35.4

FAA APPROVED


HOVER POWER +15% TORQUE (not to exceed limits) ENGINE RPM 100% GENERATOR 150 AMPS (EA.)

INITIATED FROM 1.2 METER SKID HEIGHT VTOCS = 45 KIAS HEATER ON OR OFF

4877 METERDEN.ALT.UMT

iij___i___ \ /I i O\\

i

7 _MAXIMUM

_

/

///

NJ

-

7

-

-

ROSS WEIGHT

FOIIIAKEOFF T

MIN\

OAT

/

/,

-

4J--+^X 00--,

_l - / ;

/-

/Ix' IIl

A/1II

IIIIJ I I-...t//

/

EATERON(21'C

20 OAT -C

40

60

100

Figure 4-7.

12

200

300 TA EOF

400 CE m

Takeoff distance (Metric)

600

00

BHT-412-FMS-37.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O This supplementshall be attachedto Model412EPFlight Manual when the 412-706-024Dual Flight Director (3Axis) kit has been installed.

Information contained herein supplements information of the basic Flight Manaual.For Limitations, Procedures, and PerformanceDatanot contained in this supplement, consult basic Flight Manual.

COPYRIGHT NOTICE COPYRIGHT 1999 BELL HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON INC. A DIVISION OF TEXTRON CANADA LTD. ALL RIGHTS RESERVED

Bell Helicopter


BHT-412-FMS-37.4

NOTICEPAGE


Data portion of this supplement is Manufacturer's proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron, Inc.


NP


LOG OF REVISIONS

Oct 94 0......................18 Original ....................... Revision ......................1..................... 20 Dec 99

LOG OF PAGES

REVISION NO.

PAGE

PAGE

REVISION NO.

FLIGHT MANUAL

MANUFACTURER'S DATA

Title ............................................................1 NP ..............................................................0 A -B ........................................................1 0 ii ............................................................... 1 ................................................................. 0 2 - 3 .........................................................1 4 .................................................................0 5 - 6 .........................................................1

7 - 12.......................................................0

NOTE


A


FAA APPROVED


Original ....................... 0.....................18 Oct 94 Revision...................... 1 .................... 20 Dec 99

APPROVED

MANAGER


Rev. 1

DATE


FAA APPROVED

GENERAL INFORMATION

The Dual Flight Director (3-Axis) assists the flight crew in control and navigation of the helicopter. Nine flight director modes of operation, in addition to standby (SBY) mode, are available for selection by the flight crew: Altitude (ALT), Indicated Airspeed (IAS), Vertical Speed (VS), Heading (HDG), Navigation (NAV), Instrument Landing System with Localizer and Glideslope (ILS), Back Course (BC), VOR Approach (VOR APR), and Go-Around (GA). Flight director operation Is controlled by the mode selectors on the instrument panel.

The Dual Flight Director (3-Axis) can be coupled to the Automatic Flight Control System (AFCS) for fully automatic (hands off) flight path control.

ABBREVIATIONSAND ACRONYMS ADD

-

Air Data Display (AL-300)

ALT

-

Altitude

AP1/AP2 BC DCPL TRIM/FD GA HDG HSI

-

IAS ILS

-

NAV SBY VOR

-

VOR APR VS

-

Autopilot 1/ Autopilot 2 Back Course Decouple TRIM/Flight Director Select Go-Around Heading Horizontal Situation Indicator Indicated Airspeed Instrument Landing System with Localizer and Glideslope Navigation Standby Very High Frequency Omnidirectional and Radio Range VOR Approach Vertical Speed

BHT-412-FMS-37.4


1-1. INTRODUCTION

Limitations specified in BHT-412-FM-4 and as applicable remain supplements modified herein. Operation of the Dual Flight Director (3-Axis) does not relieve the to monitor flight crew of responsibility instruments and comply with all helicopter limitations.


The Dual Flight Director (3-Axis) may be VFR or IFR nonicing used during conditions.

1 test (level AFCS preflight Digital minimum) shall be accomplished prior to first flight of the day or before planned flight into IMC.

1-6. WEIGHTAND CENTEROF GRAVITY Actual weight change shall be determined and ballast after kits are installed readjusted, if necessary, to return empty weight CG within allowable limits.

1-7. AIRSPEED

Minimum airspeed for coupled operation of flight director is 60 KIAS.

1-5. CONFIGURATION

1-22. FLIGHTOPERATIONS

(3-Axis) Kit The Dual Flight Director system, includes a modified pitot-static two mode selectors, two air data sensors, parts. The flight and miscellaneous director is operated in conjunction with two flight control computers. The Copilot Instrument Kit and the Radar Altimeter Kit must be installed.

1-22-A. NORMALVFR/IFR

Flight director coupling to autopilot is not possible with only one autopilot engaged.

1-5-A. REQUIREDEQUIPMENT 1-5-A-1. AFCS

1-5-A-2. REQUIRED EQUIPMENTIFR 1

T N Section 2 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-37.4

2-1.

INTRODUCTION

Normal procedures specified in BHT-412-FM-4 as remain applicable and supplements modified herein.

2-3.

PREFLIGHT CHECK

2-3-B.

EXTERIOR CHECK

3. FUSELAGE -

AFT LEFT SIDE

Flight control computers - Secured; compartment door secured.

2-6.

2-6J.

SYSTEMS CHECK AFCS CHECK

NOTE AFCSAOG conditions satisfied when collectivefull down, radio altimeterless than 15feet, and airspeed less than 60 knots.

After completion of level 1 AFCS check In Basic Flight Manual (BHT-412-FM-4),press NAVSTBYswitch on cyclic grip or press SBY on mode selector.Continuewith level 2 AFCS check.

2

light to extinguish.

Perform lamp and button test -

Verify all buttons on both mode selectors flashing.

Verify SAS/ATT and CPL buttons on autopilot controller flashing.

Press each flashing button individually to extinguish.

AL-300 Air Data Display - Verify END displayed. Autopilot Controller - Verify TEST button flashing. (Indicates no error codes present.)

AL-300 Air Data Display - If error code is displayed: Press SBY button on mode selector once or press NAV STBY switch on cyclic stick once to view next error code. Repeat step until END is displayed.

For information on Dual Flight Director (3-Axis) error codes, refer to table 1-1 in Manufacturer's Data. Also, for information on digital AFCS error codes, refer to Basic Flight Manual (BHT-412-FM-4).

AL-300Air DataDisplay - Verify LE-2 displayed.

Autopilot Controller - Press flashing TEST button to exit level 2 test. Air Data Display reverts to dashes. Press AP1 and AP2 to engage autopilots.

AP1,AP2, DCPL,AUTOTRIM,and MASTERCAUTIONAnnunciatorsVerifyflashing. PressMASTERCAUTION

Autopilot Controller - Press TRIM/FD to select desired Autopilot or Flight Director.

Rev. 1

BHT-412-FMS-37.4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O CAUTION

CAUTION

GROUND PROLONGED FOR BE SHALL OPERATION, AFCS OPERATED IN SAS MODE ONLY. IF AFCS IS LEFT ENGAGED IN ATT MODE DURING GROUND OPERATION, IT CAN DRIVE THE CYCLIC STICK TO A CONTROL STOP.

BEFORE TAKEOFF

2-7.

2-7-C.

DUAL FLIGHT DIRECTOR

Autopilot Controller - Press TRIM/FD 1 or 2 to select Flight Director 1 or 2 as desired.

Mode Selector - When Flight Director 2 is selected, Mode Selector for pilot is activated. When Flight Director 1 is selected, Mode Selector for copilot is activated.

2-8.

TAKEOFF

Other - Observe limitations In Basic Flight Manual (BHT-412-FM-4) and applicable Flight Manual Supplements.


Flight Director selected.

Verify Flight Director

Flight Director - After reaching 60 KIAS select modes as desired.

(AHRS Equipped and Subsequent)

Helicopters -

2-10.

2-10-A.

DESCENT AND LANDING APPROACH

Flight Director - Select mode for type of approach to be flown. Collective pitch - Adjust to maintain desired approach speed.

2-10-B.

GO-AROUND

Collective/FD mode select panel GOAROUND button - Depress at Missed Approach Point.

Airspeed - Observe limitation.

2-9.

WHEN AHRS IS SWITCHED FROM DG WILL IT MODE, MAG TO INSTANTANEOUSLY SYNCHRONIZE TO DETECTED MAGNETIC HEADING. HEADING AVOID INVALID TO SYNCHRONIZATION DUE TO FLUX VALVE SENSING ERROR DURING TURNS, MAKE SWITCH SELECTION ONLY IN LEVEL FLIGHT AND FLIGHT DIRECTOR UNCOUPLED (STBY).

S/N 36248

Collective pitch - Adjust to desired climb power setting.

Airspeed - Adjust collective setting during climb to achieve 70 KIAS and 750 fpm climb.

Mode selector - Upon completion of goaround climb, select modes as desired.

2-10-C.

LANDING

Cyclic NAV STBY button or SBY button on coupled mode selector - Depress at or above 60 KIAS to decouple selected flight director.

Rev. 1

3


3-1.

FAA APPROVED

INTRODUCTION

Emergency/malfunction procedures specified in BHT-412-FM-4 remain applicable as modified herein.

3-9. AUTOMATIC FLIGHT CONTROLS SYSTEM

3-13. STATIC PORT OBSTRUCTION

In the event of an apparent malfunction of the pilot altimeter, vertical speed indicator proceed as indicator, and/or airspeed follows: STATIC SOURCE valve -

ALTN.

Coupled flight director vertical modes Use at pilot discretion.

3-9-F. SELECTEDFLIGHT DIRECTORFAILS TO COUPLE

3-15. CAUTION AND ADVISORY LIGHTS

PROCEDURE:

Caution lights related to the Dual Flight Director are described in table 3-1. Warning flags are described in table 3-2.

AP1 AND AP2 buttons SAS/ATT button CPL button illuminated.

ON.

ATT.

Depress if CPL light not

NOTE

Both AP1 and AP2 must be operative before selected flight director will couple.


BHT-412-FMS-37.4

Table 3-1. Caution lights (amber)

PANEL

WORDING

FAULT CONDITION

CORRECTIVE ACTION

DCPL


Ensure that AP1 and AP2 are engaged in ATT mode. Engage flight director modes as desired. Depress CPL button on autopilot controller if CPL light not illuminated.

Table 3-2. Warning flags

FLAG FLAG WORDING LOCATION

FAULT CONDITION

CORRECTIVE ACTION

ATT

Vertical gyro not turning at full speed/attitude information unreliable. Associated autopilot will disengage, flight director will decouple.

If only ATT flag is displayed, pilot shall monitor standby attitude indicator.

ADI

Check PILOT and CPLT ATTD SYS circuit breakers - in. If flag does not retract, continue flight in ATT or SAS mode.

FD

ADI

Flight director not coupled due to flight director failure. (Pitch and roll command bars may or may not retract from view.)

Check FLT DIR (AC and DC) circuit breakers in. Check desired mode(s) engaged. If flag does not retract from view, continue flight in ATT or SAS mode. (ATT mode for IMC conditions.)

GS (No legend on HSI)

ADI and HSI

Glideslope signal unreliable. (ADI pitch command bar retracts from view. AFCS holds pitch attitude present at time of signal failure.)

Continue flight in any mode, using cyclic ATTD TRIM switch and collective to maintain airspeed and glidepath. (Monitor raw glideslope data.)

HSI

NAV(No legend)

VOR or localizer signal unreliable. (ADI roll command bar retracts from view. HSI course deviation bar and No. 1 bearing pointer unreliable. AFCS holds roll level attitude.)

Check NAV 1 (DC) circuit breaker in. Check VHF NAV 1 tuned properly, signal identified.Continue flight in HDG or ATT mode.

Rev. 1

5

T N I R P N O R I O T F C U T D O Section O 4 N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-37.4

Table 3-2. Warning flags

FLAG FLAG WORDING LOCATION HSI

OFF

FAULT CONDITION

CORRECTIVE ACTION

Directional gyro failure or HSI failure.Heading information unreliable. (AFCS holds roll level attitude.)

Check PILOT and CPLT HSI and GYRO CMPS (AC) circuit breakers in. Continue flight In any mode except HDG.

AHRS EQUIPPED HELICOPTERS S/N 36248 AND SUBSEQUENT

FD ATT RT ADI

ADI failure.

Check Pilot and Copilot 115 Vac ATT SYST circuit breaker.

GS ATT RT ADI

AHRS failure.

Check Pilot and Copilot 26 Vac AHRS circuit breaker.


6

Rev. 1

T N I Section 5 R P N O R I O T F C U T D 1 O Section O N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-37.4

MANUFACTURER'S DATA


1-1. INTRODUCTION

The Dual Flight Director (3-Axis) is designed for use as a workload reliever to assist the pilot or copilot in control and navigation of the helicopter. The flight director has nine modes of operation, any of which may be coupled to the autopilots for fully automatic hands off flight path control. When decoupled from the autopilots, the flight director provides automatic flight path computation and visual pitch, roll, and yaw command indications to direct the pilot or copilot in maneuvering the helicopter to maintain the selected flight path. When the flight director is coupled in the appropriate modes, the automatic flight control system will maneuver the helicopter to perform the following functions: Maintain a constant pressure altitude (ALT).

Maintain a constant indicated airspeed (IAS). Maintain a constant vertical speed climb or descent (VS).

Turn to and maintain magnetic heading (HDG).

a selected

Capture and track a selected VOR radial (NAV or VOR APR).

Capture and track an ILS localizer and glideslope (ILS). Capture and track a localizer course (BC).

back

Initiate a missed approach (go-around) climbout (GA).

The flight control computer analyzes vertical and lateral flight and navigational data to generate pitch and roll steering commands that are displayed visually on the attitude director indicator (ADI). The vertical channel combines pitch attitude, airspeed, altitude, vertical speed, and glideslope deviation information to produce computed pitch command signals. The lateral channel combines roll attitude, heading, and course deviation information to produce computed roll command signals. Automatic flight path control is achieved when the pitch and roll commands from the flight director

7


computer are coupled to the flight control computers.

Should a flight or navigation data signal become invalid, the affected pitch or roll channel will revert to the attitude hold mode, and the respective command bar on the ADI will retract from view. If either autopilot fails or is disengaged, the flight director will decouple automatically.

For further information, including complete details and specifications, refer to the Honeywell SPZ-7600 Integrated Flight Control System Pilot's Manual for the Bell 412.

1-22. FLIGHT DIRECTOR CONTROLS 1-22-A. MODESELECTOR

The mode selector (figure 1-1) enables the pilot or copilot to select the desired flight director mode by depressing the appropriate push-on/push-off button. Selected modes are annunciated by illumination of the respective mode select buttons. The NAV, BC, and VOR APR buttons have two lights each (ARM and CAP) to advise the pilot and copilot of the status of the flight director computer. The ILS button legends (ARM and GS) advise when the computer has armed or captured the glideslope. Depressing the SBY button disengages all modes, and tests the illumination of all mode selector lights. Turning the PILOT INSTR LT knob clockwise dims all mode selector lights for night operation.

1-22-B. COUPLEBUTTON

The couple (CPL) button is a push-on/ push-off button located on the autopilot controller on the pedestal (figure 1-2). When both autopilots (AP1 and AP2) are engaged in the attitude retention (ATT) mode, selecting any valid flight director mode will automatically couple the flight director to the autopilots, as indicated by

MANUFACTURER'S DATA

illumination of the ON legend of the CPL button. The pilot or copilot may decouple the flight director by depressing the CPL button. When decoupled, the flight director will continue functioning in the selected mode, providing visual pitch and roll commands on the attitude flight director (ADI). Once depressed, the CPL button must be depressed again to recouple any flight director mode.

1-22-C. FORCETRIMSWITCH

The FORCE TRIM switch, located on the miscellaneous control panel on the pedestal, must be ON anytime the flight director is coupled for automatic flight control.

The cyclic mounted FORCE TRIM release button can be depressed to allow the pilot to reposition the cyclic control and pedals manually for large scale pitch, roll, and yaw corrections. Upon depressing the button, the rotary trim actuators are deenergized, the flight director modes are decoupled momentarily, the autopilot pitch, roll, and yaw linear actuators return to centerpositions, and the autopilot computers are placed in a fast follow-up mode to track flight control positions. Upon releasing the button, the autopilots and flight director will resume functioning in the preselected modes. If previously decoupled in the ATT mode, the autopilots will maintain the attitude existing at the time the button is released.

1-22-D. ATTITUDETRIM SWITCH

The cyclic mounted attitude trim (ATTD TRIM) switch can be moved fore and aft to adjust pitch attitude and laterally to adjust roll attitude during decoupled operation in the ATT mode. Also, the ATTD TRIM switch is used to make small pitch attitude changes when coupled in any mode except ALT and ILS after glideslope capture. Roll attitude can also be adjusted by means of the ATTD TRIM switch, except when a lateral mode is engaged. (Large attitude


BHT-412-FMS-37.4

changes should be made by depressing the cyclic FORCE TRIM release button.)

for 5 seconds following mode engagement, beep trimming, or force trim release. After 5 seconds, the display reverts to dashes.

1-22-E. STANDBY BUTTON

To indicate fault conditions, various error codes are displayed on the AL-300 Air Data Display. Refer to table 1-1 for a list of error codes applicable to Dual Flight Director (3-Axis). Also, for information on digital AFCS error codes, refer to Basic Flight Manual (BHT-412-FM-4).

The cyclic mounted NAV STBY button is a remote switch having the same function as the standby (SBY) button on the mode selector (figure 1-1). Depressing the button disengages all flight director modes, tests the illumination of all mode selector lights, retracts the pitch and roll command bars on the ADI, and places the flight director in the standby status. The pilot or copilot must then reselect the modes if continued flight director operation is desired.

1-22-F. GO-AROUND BUTTON

The GO-AROUND button, located on the collective control heads (figure 1-2), is a remote switch having the same function as the go-around (GA) button on the mode selector (figure 1-1). Depressing either of the four buttons places the flight director in the go-around mode and disengages all other modes. In the go-around mode the flight director commands a roll level attitude and a pitch attitude that provide a 750 foot-per-minute rate of climb. (The pilot or copilot must adjust collective pitch to maintain desired airspeed.)

NOTE

Some error codes may not be applicable depending on equipment installed.

1-24. AUTOPILOT CONTROLLER

Either autopilot system can be operated in the SAS (Stability Augmentation System) mode or the ATT (Attitude Retention) mode. The SAS mode provides short term stabilization for hands-on flying. The ATT mode is used for hands-off flight when the autopilots are coupled to the flight director. The PC-700 autopilot controller (figure 1-4) is used to select the autopilot, SAS or ATT mode, test, trim/flight director, and coupling.

1-25. INSTRUMENTPANEL AND CONSOLES

All caution and advisory lights are dimmed using DIM switch on caution panel.

1-23. AL-300 AIR DATA DISPLAY

The AL-300 Air Data Display (figure 1-3) presents a digital readout of vertical velocity in feet per minute or indicated airspeed (IAS) in knots. Display will appear

9


MANUFACTURER'SDATA

Figure 1-1. MS-700 Mode Selector

COPILOT SIDE

412-FSMD37.4-1-2

Figure 1-2. Collective control heads


MANUFACTURER'S DATA

412-FSMD37 4-1-3

Figure 1-3.

Figure 1-4.

AL-300 Air Data Display

PC-700 Autopilot Controller

T N I R P N O R I O T F C U T O OD N R E P E Section 2 E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-37.4

MANUFACTURER'S DATA

Table 1-1. Dual flight director (3-axis) error codes

PWR UP

PRE FLT

CONT TEST

AL-300 ADD

NOTE

Handling and procedures for Director (3-Axis)

servicing Dual Flight are found in


applicable maintenance manuals provided by the equipment manufacturer.

T N I * INeDll'1aS^ R P N O R * ROTORCRAFT I O T FLIGHT MANUAL F C U T 09 O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-38.4

i

MODEL

~SUPPLEMENT

EFIS WITH FOUR-AXIS AFCS KITS 412-705-009 AND 412-705-024 WITH OR WITHOUT INCREASED MC POWER KIT 412-706-029 BHT S/N 36087 AND SUBSEQUENT CERTIFIED 18 OCTOBER 1994

This supplement shall be attached to Model 412EP Flight Manual when Electronic Flight Instrument System (EFIS) with Four-Axis AFCS Kits with or without Increased Maximum Continuous Power Kit is installed.

O


COPYRIGHT NOTICE 2000_ COPYRIGHT BELL D HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON INC. A DIVISION OF TEXTRON CANADA LTD. ALL RIGHTS RESERVED

Bell HlIicoptSer

l[I

.I

A Sulidiay of Textron ec. PT

OFICE

AOX 412

REISSUED -

FOT WOmTH. 1a

71101

29 JANUARY 1999

REVISION 2 -16

AUGUST 2000


NOTICEPAGE




BHT-412-FMS-38.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O ROTORCRAFT

FLIGHT

MANUAL

SUPPLEMENT

EFIS WITH FOUR-AXIS AFCS KITS 412-705-009 AND 412-705-024 WITH OR WITHOUT INCREASED MC POWER KIT 412-706-029 TEMPORARY REVISION FOR DUAL GENERATOR FAILURE OPERATIONS S/N 36218,36220,36222,36227, AND 36233 CERTIFIED 19 OCTOBER 1994

tnsert these Temporary Revision numbered pages in supplement.

pages next to like

DO NOT remove existing pages from supplement.

information contained herein supplements information of Basic plight Manual. For limitations, Procedures, and performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helkopter~

&ND BELLHEUCOPTERTEXTFON INC. ~Lyl~~o~~FmmR~ C&MDA LTD.

REISSUE29 JANUARY 1999 TEMPORARY REVISION - 02 JULY 1999 REISSUE - 21 JANUARY 2000

BHT-412-FMS-38.4

T N MODELR RI P N ROTORCRAFT O R I FLIGHT MANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SUPPLEMENT EFIS WITH FOUR-AXIS AFCS KITS 412-705-009 AND 412-705-024 WITH OR WITHOUT INCREASED MC POWER KIT 412-706-029 TEMPORARY REVISION FOR INVALID VOR DATA DISPLAYED ON EFIS S/N 36087 - 36247

CERTIFIED 18 OCTOBER 1994

Insert these Temporary Revision pages next to like numbered pages in supplement. This T.R. remains active until compliance with ASB 412-00-104. DO NOT remove existing pages from supplement.

Information contained herein supplements information of Basic Flight Manual. For limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell HelicopterCEai

COPYRIGHT NOTICE 2000 COPYRIGHT BELL HELICOPTER TEXTRON INC. INC.C TEXTRON ANDBELLHELICOPTER A DIVISIONOF TEXTROCANADA LTD. ALL RIGHTS RESERVED

O

.o

.Ex

oWORT

A , . 1EJI ,a11

C

A

REISSUE 29 JANUARY 1999

TEMPORARY REVISION -

REISSUE 16 AUGUST 2000

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-41%FMB-38.4

LOGOFREVISIONS

Temporary Temporary

18 DEC 98 Reissue ... . ...... 0, FEB 99

Temporary Reissue .......... .. 02 WI. 99 Temporary . Reissue 30 BEP 99 Temporaty ......... .... Reissue 21 JAN 00

LOGOFPAGES REVISION

PAGE

NO.

PAGE

REVISION NO.

FLIGHT MANUAL

Title ..................................... .... ......... Temporary A - B ______._._........,.,.,,,,.,,,,.,.,,.,......... Temporary 7 ........................... . . .. . . . ........... Temporary

NOTE

Revised text is Indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. TEMPORARY REVlSlON - 02 WLY ,998 REISSUE - 21 JANUARY 2ooO

A


LOG OF REVISIONS

Temporary ......................................... 26 May 00 Temporary ............ Reissue.............. 16 Aug 00

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title.................................. Temporary Reissue A- B .............................. Temporary Reissue

REVISION NO. PAGE Reissue Temporary 10 .................................... MANUFACTURER'S DATA 21A /21B ......................... Temporary Reissue

NOTE Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. TEMPORARY REVISION -


A


LOG OF REVISIONS

Original ....................... 0......................18 Oct 94 Revision ...................... 1..................... 15 Apr 98 Reissue ....................... 0..................... 29 Jan 99

Revision ...................... 1 .................... 26 May 00 Revision ...................... 2.................... 16 Aug 00

LOG OF PAGES

PAGE

REVISION NO.

REVISION NO.

PAGE

Title............................................................2 NP .............................................................. 0 A - B ........................................................ 2 i/ii .............................................................. 2 1 - 3 ......................................................... 0 4 - 6 ......................................................... 2 7 ................................................................ 0 8 ................................................................. 2 9 ................................................................. 0

10 -11 ....................................................... 2 12--16 .................................................... 0 17/18 .......................................................... 0 19/20 . . ...................................... 0 MANUFACTURER'S DATA 21 - 23 ..................................................... 24 ............................................................... 25 - 30 ..................................................... 31 -33 ..................................................... 34 ...............................................................

0 2 0 2 0

NOTE


16 AUG 00

A

FA&APPROVED

BHT-412.FMS-38.4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOGOFFAAAPPROVED REVISIONS Temporary . . . . . . 18DE098 Temporary Reissue 02JUL99 _..._._._..__ Reissue _____.__._..__ 30SEP99 Temporary Reissue 01FES99 Temporary Temporary _._._..____._ Reissue ____.__._._._ 21JAN00

NOTICE

ThisTemporary Revision expires 31JUL2000.

DATE

ROTORCRAFT CERTIFICAT,ON OFFICE FEDERAL AWATION ADMlNlSTRATlON FT.WORTH. TX76193-0170 B TEMPORARY REWSION - 02JULY 1998 REISSUE - 21JANUARY 2000


BHT-412-FMS-38.4


Temporary ........................................ 26 May 00

Temporary ............ Reissue .............. 16 Aug 00

APPROVED

#

4A

/

t

*MANAGER


TEMPORARY REVISION - REISSUE 16 AUGUST 2000

DATE

A4

dS/

//,

2&&&

T N I R P N O R I 0 O T F C U T O OD N R E P E E R R C S N G O N I R W O E 0i I F Y V NL O BHT-412-FMS-38.4

FAA APPROVED


Original ........................ 0 Reissue ........................ 0 Revision.......................1 Revision.......................2

..................... 18 Oct 94 ..................... 29 Jan 99 .................... 26 May 00 .................... 16 Aug 00

APPROVEDED#

-/-t

'MANAGER


Rev.2

16AUG00

DATE

/

2,zs


FAA APPROVED

GENERAL INFORMATION

The 412EP Electronic Flight Instrument System (EFIS) with Four-Axis AFCS kits with or without Increased Maximum Continuous Power kit configuration includes the Honeywell SPZ-7600 Integrated Flight Control System. This system consists of DFZ-760 Digital AFCS, EDZ-705 EFIS, and AA-300 Radar Altimeters. Refer to the appropriate manufacturer's publication for operations. ABBREVIATIONS, ACRONYMS, AND PLACARDING b PLACARDING

ESNTL

Abbreviations, acronyms, and placarding used throughout this manual are defined as follows:

-

Essential

FD

-

Flight Director

HDG HSI

-

Heading Horizontal Situation Indicator

ADD

-

Air Data Display (AL-300)

IAS ICS

-

Indicated Airspeed Intercommunication system

ADF

-

Automatic direction finder

IDENT

-

Identification

AHRS

-

ILS LE

-

Instrument Landing System Level

ALT

-

Attitude Heading Reference System *System Altimeter

AP1/AP2

-

Autopilot 1/Autopilot 2

LF LH LH

Low Frequency Left Hand Left Hand

ATT or ATTID

-

Attitude

-

NAV

-

Navigation

-

Gas producer RPM

PRYC

-

Pitch/Roll/Yaw/Collective

BLWR BC

-

Blower Back Course

NG

CBOX

-

Combining Gearbox

PWR

-

Power

CDI

-

Course Deviation Indicator

RADALT

-

Radar Altimeter

CLTV CPL

-

Collective Couple

REL REV

-

Release Revert

COLL

-

Collective

RH

-

Right Hand

CPLT

-

Copilot

STBY

-

Standby

DCPL

DECEL DEG

-

Decouple Deceleration Degrees

SYS VAC VDC

-

System Volts Alternating Current

-

Volts Direct Current

DH

-

Decision Height

VEL HLD

-

Velocity Hold

bDME

-

Very High Frequency VHF omnidirectional range

-

VHF VOR VOR APR WAT

-

EADI

-

VOR Approach Weight Altitude Temperature

EFIS

-

Distance Measuring Equipment EFIS Attitude Director Indicator Electronic Flight Instrument

WPT

-

Waypoint

EHSI

-

EFIS Horizontal Situation Indicator

WX XPDR

-

Weather Transponder

ENC

-

Encoding

W^y~

~System

Rev. 2

16 AUG 00

i/ii

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-38.4

FAA APPROVED

1-1. INTRODUCTION

Limitations specified in BHT-412-FM-4and supplements remain applicable. Operation of SPZ-7600system does not relieve flight to monitor crew of responsibility instruments and comply with all helicopter limitations.

During all flight operations, Honeywell SPZ-7600 Integrated Flight Control System Pilot's Manual for Bell 412 will be accessible to the pilot.

1-5. CONFIGURATION 1-5-A.

When referring to required equipment lists in BHT-412-FM-4,note that EFIS replaces ADI/HSI.

1-6. WEIGHT AND CENTER OF

GRAVITY

Actual weight change shall be determined after kits are installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits.

For ground operations when OAT is between 48.9 and 51.7°C (120 and 125°F) with doors and windows closed and EFIS fans operating, EFIS operation is limited to 90 minutes.

1-13. POWER PLANT Applicable engines.

to PT6T-3D, 3DE, or 3DF

WARNING

THE FZ-706 FLIGHT CONTROL DOES NOT COMPUTER RECOGNIZE ENGINE FAILURE. FOLLOWING ENGINE FAILURE, COMPUTER WILL CONTINUE TO LIMIT REMAINING ENGINE TO THE APPROPRIATELIMITS UNTIL PILOT DISENGAGESCOLLECTIVE COUPLING OR MANUALLY COLLECTIVE OVERRIDES LIMITING. UNDER CERTAIN LOW OAT CONDITIONS THE AFCS COMPUTERMAY EXCEEDSINGLE ENGINETORQUE LIMIT.

1-8. ALTITUDE

1-13-C. INTERTURBINE TEMPERATURE

Maximum operating pressure altitude is 20,000 feet - collective coupling is limited to 16,000 feet pressure altitude.

Engines are limited to 810° ITT during all coupled modes of operation.

1-11. AMBIENTTEMPERA-

1-14. TRANSMISSION

TURES

The Honeywell DFZ-7600 system, in the 4axis configuration, can be set to recognize 1


BHT-412-FMS-38.4

or the either the basic transmission Maximum Continuous Power Kit (412-706-

029).

Depending upon the configuration, are as follows:

limits

Transmission Is limited to 100% MAST TORQUE (or 810 ° engine ITT) when airspeed is below 60 KIAS (or in GOAROUND mode with airspeed below 105 KIAS). When airspeed is above 60 KIAS, transmission is limited to 81% MAST TORQUE. Maximum 706-029):

Continuous

TOLERANCES MAY ALLOW BASIC HELICOPTER LIMITATIONS TO BE EXCEEDED FOR SHORT DURATIONS. IT REMAINS THE RESPONSIBILITY OF THE PILOT TO MONITOR AND CORRECT FOR ANY EXCEEDANCES.

Flight director shall not be coupled autopilots below 60 KIAS.

to

Flight director coupling to autopilot is not possible with only one autopilot engaged. Uncoupled ILS approaches shall flown in 3-cue DECEL mode.

not be

Power Kit (412-

CAUTION

Transmission is limited to 81% MAST ° TORQUE (or 810 engine ITT) when airspeed is above 105 KIAS and 100% is MAST TORQUE when airspeed below 105 KIAS.

1-22. FLIGHTOPERATIONS

1-22-A. EFISCOMPOSITEMODE

The EFIS composite mode is to be used in IFR flight only after in-flight failure of an EADI or EHSI. Use of composite mode for VFR flight is allowed.

1-22-B. NORMALVFR/IFR

test (level 1 AFCS preflight Digital minimum) shall be accomplished prior to first flight of the day or before planned flight into IMC. WARNING

DURING NORMAL AFCS COUPLED OPERATION, SYSTEM

THE RADALT MODE IS NOT INTENDED TO FUNCTION IN A TERRAIN FOLLOWING OR TERRAIN AVOIDANCE MODE IN FORWARD FLIGHT. THE COLLECTIVE COMMAND IS RATE LIMITED, AND WILL NOT CAUSE THE HELICOPTER TO CLIMB OVER RAPIDLY RISING TERRAIN.

hold of velocity For proper operation (VELHLD) mode or radar altitude (RADALT) mode, as activated from the autopilot controller, refer to Honeywell SPZ-7600 Integrated Flight Control System Pilot's Manual for the Bell 412.


FAA APPROVED

2-1. INTRODUCTION

2-6. SYSTEMSCHECK

Normal procedures specified in BHT-412FM-4 and supplements remain applicable as modified and added to in this section. Operation of SPZ-7600 system does not relieve flight crew of responsibility to monitor and comply with all helicopter normal procedures.

2-6-B. FORCETRIMCHECK

2-3. PREFLIGHTCHECK 2-3-B. EXTERIORCHECK 3. FUSELAGE-

AFT LEFT SIDE

Secured; AFCS computers compartment door secured.

Collective control head switches All LT rheostats -

OFF.

OFF.

OFF.

EFIS MASTERswitch EFIS control panel -

FORCETRIM switch - ON, cover down. CLTV/YAW TRIM switch - OFF. Check and pedal force trim collective disengage. CLTV/YAWTRIM switch - ON.

Flight controls - Move to check PRYC out of detent caution lights illuminate (P = pitch,

R = roll,

collective).

Y = yaw, and C =

2-6-J. AFCSCHECK

2-4. INTERIORAND PRESTARTCHECK CLTV/YAWTRIM switch -

Flight controls - Ensure cyclic and collective friction set to minimum. Collective lock removed.

OFF.

As desired.

NOTE Verification of AFCS actuator centering is necessary. Failure of the actuators to center could result in reduced control margins (pitch and roll) and abnormal control positions (pitch, roll, and yaw). EFIS MASTER switch - ON. EFIS control panel - As desired.

Pilot and copilot EFIS attitude indicators - Erect and set intensity as necessary.

2-5. ENGINESTART 2-5-A.

Collective pitch RPMincreases.

Lower as ROTOR

CLTV/YAWTRIM switch - ON.

WARNING


3


BHT-412-FMS-38.4

Press SBY to continue with level 2 test if desired, or press TEST switch to exit AFCS test.

AFCS system - Test as follows:

Verify AP1 and AP2 are disengaged. Cyclic and collective force trim switches - Check ON.

2-6-K.

AFCS TEST switch - ON for level 1 test.

Perform AHRS test as follows:

Verify AL-300 Air Data Display reads Lg 1.

AHRS CHECK (S/N 36248 AND SUB)

AHRS Control Panel TEST switch No. 1 AHRS Control Panel TEST switch No. 1 and No. 2 - Press and hold.

w

Verify AP1 light flashes, then press.

NOTE

NOTE

Pilot shall closely monitor controls during test, taking care not to restrict To prevent control movements. nuisance failures, hands and feet must remain off controls.

EHSI


Heading HDG Flag HDG Flag

015 ° In view In view

Pedals - Verify aft and forward movement (left pedal then right) approximately one-half inch.

EADI ATT Flag RT Flag

In view In view


AHRS CONTROL PANEL


in DG or MAG in DG mode, 000° at test realigned.

Verify following responses for each system:

Collective - Verify movement up then down approximately one inch.


4

Test may be performed mode. If accomplished heading will remain completion and must be

/ +

2-7.

Needle deflection full -

BEFORE TAKEOFF

Engine, gearbox, transmission,

Pedals - Verify aft and forward movement (left pedal then right) approximately one-half inch.

hydraulic, and electrical instruments Within operating ranges. Caution and warning lights -

Upon completion of test, verify AP1 and AP2 lights extinguish and AL-300 Air Data Display reads END 1.

Extinguished. Flight instruments - Check operation and set. Select required EFIS display.

If error codes are present, press SBY (or cyclic NAV STBY switch) once to view each code until END 1 is displayed (refer to Section 3 of this supplement for list of error codes).

When flying IFR check as follows: IVSI - Needle near zero.

Rev. 2

16 AUG00

Altimeter - Set and check. EHSI - Check.

0


BHT-412-FMS-38.4

EADI - Check.

Radar altimeter - Zero altitude; OFF flag retracted; DH set as desired. NOTE

When AFCS is in ATT mode, the cyclic FORCE TRIM release and CLTV YAW TRIM REL switches should be pressed before liftoff (to trim actuators to center positions) and should be held until desired climbout attitude is attained,

2-8.

TAKEOFF

ENG - 100% RPM (Nil).

AP1 and AP2 - ON.

selected in level flight to avoid heading and navigational errors. and errors 2-9-B. MANEUVERING WITH AFCS IN ATT MODE Press cyclic FORCE TRIM release and CLTV YAW TRIM REL switches and maneuver as desired. Release switches when desired attitude is reached. Autopilot will hold attitude until retrimmed to new attitude. Attitude may also be adjusted with cyclic ATTD TRIM switch.

For momentary attitude changes, manual cyclic movement may be used; however, AFCS actuators may be saturated to limit authority when cyclic is moved manually.

FD1 or FD2 - As desired.

2-9.

2-9-A.

IN-FLIGHT OPERATIONS FLIGHT DIRECTOR OPEIRATIONROS 29A

Flight Director - Verify Flight Director selected.

Flight Director - After reaching 60 KIAS select modes as desired. NOTE

To perform any course change, other than at station passage, during NAV VOR APR, ILS, or BC modes while either ARM or CAP, temporarily deselect the mode. Reset the HSI for the new course and intercept heading then re-arm for capture.

Rotation of the course knob during NAV, VOR APR, ILS, or BC modes while either ARM or CAP, outside the cone of confusion can falsely trigger the over station sensor resulting in degraded performance. (For S/N 36248 and Subsequent) DG mode to MAG mode must only be

IWAR

IN THREE CUE COUPLED CLIMB (VS) OR ALTITUDE HOLD, THE FZ-706 FLIGHT CONTROL COMPUTER WILL ALLOW COMMANDED AIRSPEED TO PRESERVE TO DECREASE COMMANDED CLIMB RATE AND ALTITUDE HOLD AS A COLLECTIVE LIMIT IS APPROACHED.

WHEN IN VEL HLD MODE, SELECTION OF OF ANY ANY FLIGHT FLIGHT DIRECTOR DIRECTOR MODES MODES (PIOT OR COPILOT) WILL DISENGAGE VEL HLD. MONITOR POWER DURING HOVER OPERATIONS. COLLECTIVE LIMIT IS 100% MAST TORQUE OR 810°C ITT. MANUALLY ADJUST COLLECTIVE WITH COLLECTIVE TRIM RELEASED AS REQUIRED TO MAINTAIN POWER WITHIN TRANSITORY LIMITS. DO NOT PUSH AGAINST COLLECTIVE TRIM OR A SLOW COLLECTIVE OSCILLATION MAY DEVELOP.

IF DIGITAL AFCS COMPUTER DOES NOT RECEIVE RADAR ALTITUDE SIGNAL WHEN IN A COUPLED ILS Rev. 2

16AUG00

5


BHT-412-FMS-38.4

NOT WILL AUTO-LEVEL MODE, WILL OCCUR AND HELICOPTER CONTINUE TO FOLLOW GLIDESLOPE SIGNAL. NOTE

autopilot During single system operation (AP1 or AP2) at airspeeds above 60 KIAS, autopilot roll axis will normally hold heading. However, if excessive roll activity is experienced due to turbulence, pilot may wish to fly with yaw force trim released. This will cause autopilot roll axis to maintain lateral stability.

For helicopters S/N 36087 - 36247, inflight use of VG FAST ERECT switch will disengage the respective autopilot and decouple the automatic flight control modes.

helicopter loading certain Under conditions, smoother flight may be obtained in turbulence by selecting 2cue coupling instead of 3-cue coupling. Flight controls - Adjust (minimum friction for ATT mode operation).

WRNING

THE FZ-706 FLIGHT CONTROL COMPUTER DOES NOT RECOGNIZE FOLLOWING ENGINE FAILURE. ENGINE FAILURE, THE COMPUTER WILL CONTINUE TO LIMIT THE TO REMAINING ENGINE APPROXIMATELY NORMAL CRUISE POWER UNTIL PILOT DISENGAGES COUPLING OR COLLECTIVE MANUALLY OVERRIDES COLLECTIVE LIMITING NOTE

Irregular terrain may affect decel capture and auto-level performance.

2-10

DESCENT AND LANDING

Flight controls - Adjust friction as desired (minimum friction for ATT mode operation). low 60 60 KIAS. KIAS. FD1 or FD2 - Disengage below Veri

2-11

mode selectors - SBY.

ENGINE SHUTDOWN

AP1 and AP2 - Disengage; check autopilot lights extinguish, AFCS and MASTER CAUTION lights illuminate. EFIS MASTER switch - OFF.

6

Rev.2

16AUG00

BHT-412FMS-38


3-7.

ELECTRICAL

3-7-C.

.

3

SECOND FAILURE

SYSTEM

GENERATOR

INDICATION:

ENG 1 and ENG 2 DC GENERATOR caution llghts Illumlnafe. NOTE

Procedures outlined in 2-T-A of Basis Flight Manual will allow for 2.4 minutes of battery power after dual generator failure.

.

PROCEDURE:

Pull out the following clrcull breakers: EFfS NO 1 SLWR EFIS NO 1 SODC.

TEMPORARY REYlSlON - 02 JULY ,999 REtSSVE - 21 JANUARY 2000

7

21 Jan 2000


FAA APPROVED

I

3-1. INTRODUCTION

Emergency and malfunction procedures and in BHT-412-FM-4 specified as remain applicable supplements modified and added to in this section. Operation of SPZ-7600 system does not relieve flight crew of responsibility to monitor and comply with all helicopter emergency and malfunction procedures.

Tables 3-1 and 3-2 are provided as quick references to certain failure displays and recommended corrective action.

other electrical malfunctions. Failure of No. 1 hydraulic system will render AFCS yaw axis Inoperative but will not affect pitch or roll SAS or ATT mode functions. Failure of No. 2 hydraulic system will not affect AFCS. If both autopilots are disengaged, the following procedures do not apply.

3-9-A. AFCSFAILS TO ENGAGE OR DISENGAGES INDICATION:

AFCS caution light illuminated.

THE FZ-706 FLIGHT CONTROL NOT DOES COMPUTER RECOGNIZE ENGINE FAILURE. FOLLOWING ENGINE FAILURE, COMPUTER WILL CONTINUE TO LIMIT REMAINING ENGINE TO APPROPRIATELIMITS STATEDIN SECTION 1 - LIMITATIONS UNTIL PILOT DISENGAGESCOLLECTIVE COUPLING OR MANUALLY COLLECTIVE OVERRIDES LIMITING. UNDER CERTAIN LOW AFCS OAT CONDITIONS, COMPUTERMAY EXCEEDSINGLE ENGINETORQUELIMIT.

3-9. AUTOMATICFLIGHT CONTROLSSYSTEM

The AFCS can be affected by malfunctions of any of the vertical gyros, yaw rate gyros, inverters, accelerometers, or by

NO. 1 AUTOPILOT/NO. 2 AUTOPILOT caution light illuminated. AP1 or AP2 off (light not illuminated).

Possible erratic API indications on AP1 or AP2. Possible ATT FAIL flag displayed on pilot or copilot EADI.

Possible illumination of NO. 1 or NO. 2 INVERTERcaution light. NOTE

If inverter 1 or 2 fails, AP1 or AP2 but can be will disengage, reengaged by pressing respective switch on AFCS control panel.

PROCEDURE: Airspeed -

Reduce to 115 KIAS or less. 7


BHT-412-FMS-38.4

INV 1 and 2 switches - ON; check NO. 1 and NO. 2 INVERTER caution lights extinguished. Pilot and copilot EADIs - Check ATT FAIL flags retracted, indicators functioning properly. »-^'~ .... C CAUTION -^-....-.. N -CLTV/YAW

DO NOT ATTEMPT TO RESET ANY MORE THAN CIRCUIT BREAKER ONCE. ~~~~~~ONCE'~. ~If

Check the following circuit breakers in: INV 1 PWR and INV 2 PWR AC Feeder Panel (8)


3-9-B.

AFCS FAILS TO HOLD ATTITUDE

*PROCEDURE:

FORCE TRIM switch - Check ON. TRIM switch - ON.

SAS/ATT button - Check ATT light illuminated. AUTOTRIM caution light illuminated Follow procedure for AUTOTRIM FAILURE. FAILURE.

3-9-C.

ATTD 26VAC (1 and 2) AFCS 28 DC (1 and 2)

AFCS HARDOVER OR ABNORMAL CONTROL RMALONTROL AB DISTURBANCE

(For helicopters S/N 36087 - 36247) ATTD GYRO 115 VAC (1 and 2)

(For helicopters S/N 36248 and sub) AHRS 28 VDC (1 and 2)

AP1 OR AP2 FAILS OR IS DISENGAGED, REDUCE AIRSPEED TO

ATTD 26VAC (1 and 2)

115 KIAS OR LESS.

AP1 or AP2 switch (affected system) Press to reengage.

If either autopilot will not reengage, or if abnormal control disturbance occurs, follows: as follows: proceed as proceed

*PROCEDURE:

Cyclic FORCE TRIM release switch Press.

Affected autopilot - Disengage.

CLTV YAW TRIM REL switch - Press. Correct helicopter attitude with cyclic and pedals, then release cyclic FORCE


TRIM switch and CLTV YAW TRIM REL switch.

If fault condition cannot be cleared, use error code displayed on AL-300 to determine failed axis. Refer to table 3-3 for a list of error codes.

Airspeed - Reduce to 115 KIAS or less. Actuator position indicators - Check both systems. If any API shows maximum displacement or erratic operation of any actuator, switch affected autopilot OFF. , p , If IFR, land as soon as practical, or continue flight under VFR, if desired.

If both autopilots fail to reengage, proceed as follows: Airspeed - As desired.

8

IF

Rev.2

16AUG00


FAA APPROVED

3-9-D. AUTOTRIM FAILURE INDICATION:

AUTOTRIMcaution light illuminates.

If an attitude source fails, pitch scale, roll pointer, and aircraft symbol will be painted blue and a red ATT FAIL will be displayed in the middle of the sphere. 3-14-A-2.

PROCEDURE:


Attempt to reset AUTOTRIM light by cycling either AP1 or AP2 off, then ON.

If any of these systems fail, the respective pointer will be removed and a red X will be drawn through the respective scale. 3-14-A-3.

If fault condition cannot be cleared, use error code displayed on AL-300 Air Data Display to determine failed axis. I Refer to table 3-3 for a list of error codes. For yaw trim failures:

GLIDESLOPE, EXPANDED LOCALIZER, AND RATEOF-TURN FAILURES

INTERNAL FAILURES

Certain internal failures within the symbol generator will result in display being replaced by a red X centered on EADI with notation SG FAIL.

3-14-B. EADI CAUTIONS (AMBER ANNUNCIATIONS)

CLTV/YAWTRIM switch - OFF.

Cycle either AP1 or AP2 off, then ON.

Continue flight with manual control of collective and pedals. ATT mode may still be used, including pitch/roll FD coupled operation.

For pitch or roll trim failures:

Pilot may choose to remain in ATT mode, using cyclic FORCETRIM release switch and APIs to manually re-trim as required to keep actuators working near center; or, pilot may select SAS mode with FORCE TRIM switch ON or OFF as desired.

3-14-B-1.

SAME ATTITUDE SOURCE

If pilot and copilot have selected the same attitude source, the attitude source will be annunciated in amber. 3-14-B-2.

COMMON SYMBOL GENERATOR

When in SG REV mode, one symbol generator is driving both pilot and copilot display tubes. During this mode, a reversionary annunciation is given in amber indicating the display source. This annunciation appears in the upper left corner of the attitude sphere and will display SG1or SG2, depending on source.

3-14. EFIS

3-14-B-3.

3-14-A. EADI FAILURES (RED ANNUNCIATIONS) 3-14-A-1. ATTITUDE FAILURE

DECISION HEIGHT

When radar altitude Is within 100 feet of decision height, a white box will appear to the left of radar altitude display. When at or below decision height, an amber DHwill appear inside the white box.

9


BHT-412-FMS-38.4

3-14C-1.

HEADING FAILURES (CONT) NOTE

If amber HDG is illuminated in lower left EADI display when secondary NAV is selected, pilot shall disregard VOR course deviation and bearing pointer Information.

10


16 AUGUST 2000

Aug 16 2000


BHT-412-FMS-38.4

3-14-B-4. FLIGHT DIRECTOR FAILURE

An amber FD FAIL warning will be displayed in the upper left quadrant of EADI. Also, flight director command bars and all mode annunciations will be removed. NOTE

FD FAIL is also displayed while the digital AFCS TEST mode is engaged.

3-14-B-5. DECISION HEIGHT FAILURE

Amber dashes will replace numerical values of decision height display.

3-14-C.

3-14-C-1.

A failure in the heading system will result in removal of bearing pointers and their annunciators, course select pointer, course deviation indicator and scale, and compass sync indicator. The digital course select and heading select readouts will be replaced with amber dashes and a red HDG FAIL is displayed at the top of the compass card. 3-14-C-2.

3-14-B-7. ATTITUDE (ATT)

COMPARISON MONITOR

Amber ATT will illuminate in lower left of EADI if pilot and copilot attitude miscompare by more than 6°.

COURSE DEVIATION GLIDESLOPE DEVIATION

OR

FAILURE

3-14-B-6. RADAR ALTITUDE FAILURE

Amber dashes will replace numerical values and rising runway will be removed from display.

EHSI FAILURES (RED ANNUNCIATIONS) ( HEADING FAILURES

A failure of these systems results in removal of course or glideslope pointer, and a red X being drawn through scale. The letter G will appear at the zero deviation position of vertical scale to identify invalid information. 3-14-C-3.

INTERNAL FAILURES

Amber ATT will illuminate if cross side attitude has failed.

Certain internal failures within the symbol generator will result in display being replaced by a red X centered on EHSI with notation SG FAIL. FAIL. 3-14-D. EHSI CAUTIONS

3-14-B-8. HEADING (HDG) COMPARISON MONITOR

3-14-D-1.

Amber HDG will illuminate in lower left of EADI if pilot and copilot heading indicators miscompare by more than 60 or 120 when bank angle exceeds 6°. 6°.

NAVIGATION SOURCE Whenever pilot and copilot share same heading source or select same NAV radio, appropriate source annunciation will turn

Amber HDG will illuminate if cross side heading has fawiled.

amber. NAV radio annunciation VOR 1 or VOR also be amber if either side displays secondary navigation source.

NOTE

For monitored pFor other othser comparison comparison monitored parameters, refer to Honeywell SPZ7600 Integrated Flight Control System 412. Bell 412. for Bell Pilot's Manual for

10

Rev. 2

16AUG 00

(AMBER ANNUNCIATIONS) HEADING SOURCE AND

3-14-D-2. WAYPOINT

An amber WPT annunciation will be displayed to the left of the NAV source annunciation to indicate the passage of a long range NAV indicate the passage of a long range NAV waypoint. waypoint.

T N I R P N 0^^&b~~~~~~~ O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-38.4

Table 3-1. Caution and advisory lights

PANEL WORDING

FAULT CONDITION

CORRECTIVE ACTION

AHRS FAN

AHRS fan inoperative.

Avoid prolonged ground operations if OAT exceeds 40°C (105°F). EFIS power OFF during ground operations if OAT exceeds 35' C (95°F).

~~~Turn

DCPL


Ensure that AP1 and AP2 are engaged in ATT mode. Engage flight director modes as desired. Ensure CPL switch and COLL switch are ON.

FT OFF


Check FORCE TRIM switch ON and FORCE TRIM circuit breaker in.

AUTOTRIM

Pitch, roll, or yaw trim inoperative. Cycle either AP1 or AP2 off, then ON. If light does not extinguish, follow AUTOTRIM FAILURE procedures in this section.

CLTV

Collective autotrim inoperative.

NO. 1 AUTOPILOT/ Autopilot 1 or 2 inoperative. NO. 2 AUTOPILOT

Turn CLTV/YAW TRIM switch ON. Check excessive adjustable collective friction. Cycle either AP1 or AP2 off, then ON. If light does not extinguish, continue flight with manual collective inputs. Engage affected autopilot. If unable to engage, continue flight on remaining system.

NO. 1 EFIS FAN/NO. EFIS blower fan 1 or 2 inoperative. Pull EFIS NO. 1 or 2 circuit breakers if OAT is above 49°C (120°F). Push circuit 2 EFIS FAN in prior to landing. *i^^^~~~~~~~~ ~~~~breakers PRYC OUT OF DETENT

Pitch, roll, yaw, and collective out Release force trim of affected system of detent. and center control.

Rev. 2

16AUG00

11


BHT-412-FMS-38.4

Table 3-2. Failure displays and recommended corrective action

ANNUNCIATIONOR INDICATION

FAULT CONDITION FLIGHT DIRECTOR CORRECTIVEACTION

Red X and SG FAIL or both displays blank.

Symbol generator failure.

Flight director inoperative (no FD cues).

Press SG REV switch to revert to cross-side display and flight director.

Display reverts to power-up configuration and bearing pointers revert to NAV 1 and NAV 2.

Display controller failure.

NAV modes may disengage.

No action required. Display Controller will not operate.

Display goes blank, or color and picture quality failure.

Electronic display failure.


Turn failed display DIM knob OFF to obtain Composite Mode on remaining display.

HDG or CRS pointers Remote instrument controller failure. jitter.

None.

If only one side Is affected, SG REV may be used at pilot's discretion.

Blue sphere and ATT Attitude gyro FAIL on EADI. failure.

Command bars out of view.

Press ATT REV switch to revert to cross-side source.

Red HDG FAIL on EHSI. Loss of HDG, CRS, etc.

Directional gyro failure.

Roll command bar may bias out of view.

Press HDG REV switch to revert to cross-side source.

Red X over CDI scale; loss of bearing pointer.

NAV invalid.

None.

Check NAV tuning and station range. Select alternate NAV source if required.

12


BHT-412-FMS-38.4

Table 3-3.

PWR UP

PRE FLT

CONT TEST

X X X X X X

X X

X X

X X X X X X X X X

X X

Digital AFCS error codes AL-300


BHT-412-FMS-38.4

Table 3-3.

PWR UP

X X

PRE FLT

CONT TEST

Digital AFCS error codes AL-300 ADD

(Cont)


FAA APPROVED

Table 3-3.

PWR UP

PRE FLT

CONT TEST


X X X X X X X X X X X X X X X X X X X X X X X X X

E136 E137 E138 E139 E140 E141 E142 E143 E144 E145 E146 E147 E148 E149 E150 E151 E152 E153 E154 E155 E156 E157 E201 E202 E203

X X X

E204 E205 E206

X X

E207 E208

X

E209

X

E210

X X X X

E211 E212 E213 E214

X X X X

E215 E216 E217 E218

X

E219

X

E220

(Cont)


PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED YAW TRIM ACT MOVED YAW SERIES ACT MOVED COLL ACT MOVED PITCH TRIM (H) E/A ON PITCH TRIM (L) E/A ON ROLL TRIM (H) E/A ON ROLL TRIM (L) E/A ON YAW TRIM (H) E/A ON YAW TRIM (L) E/A ON TRIM ENABLE FAIL TRIM ONE-SHOT FAIL PITCH TRIM (H) E/A OFF PITCH TRIM (L) E/A OFF ROLL TRIM (H) E/A OFF ROLL TRIM (L) E/A OFF YAW TRIM (H) E/A OFF YAW TRIM (L) E/A OFF SERVO VALID FAIL TRIM END AROUND FAIL APON OTHER INPUT FAIL AL-300 KNOB MOTION AIRSPEED MORE THAN 60 KNOTS PRIMARY RADALT MORE THAN 20 FT ANALOG GROUND FAIL LATERAL ACCELEROMETER FAIL LONGITUDINAL ACCELEROMETER FAIL VERTICAL ACCELEROMETER FAIL ROLL ATTITUDE MORE THAN 10 DEG PITCH ATTITUDE MORE THAN 10 DEG YAW RATE GYRO MORE THAN 10 DEG/SEC DIRECTIONAL GYRO FAIL VERTICAL GYRO NOT VALID DIRECTIONAL GYRO NOT VALID PRIMARY RADAR ALTIMETER NOT VALID AIR DATA SENSOR NOT VALID SERIAL I/O NOT VALID OTHER AUTOPILOT NOT IN TEST MODE SELECTOR BUTTON(S) PRESSED AP CONTROLLER BUTTON(S) PRESSED -

15


BHT-412-FMS-38.4

Table 3-3.

PWR UP

PRE FLT

CONT TEST


(Cont)


X

E221

HARDWARE NOT VALID

X

E222

FOUR-WAY BEEP SWITCH PRESSED

X

E223

ENC ALT LESS THAN -600 FT OR GREATER THAN 12,000 FT

X

E224

BARO ALT LESS THAN -600 FT OR GREATER THAN 12,000 FT

X

E225

POD DURING TEST

X

E226

ROD DURING TEST

X

E227

YOD+YFTR DURING TEST

X

E228

COD DURING TEST

X

E229

FTR DURING TEST

X

E230

CFTR DURING TEST

X

E231

TRIM SWITCH OFF DURING TEST

X

E232

X

E233

COLL TRIM SWITCH OFF DURING TEST ENG TORQUE LESS THAN 5% OR GREATER THAN 30%

X

E234

ENG TEMP LESS THAN 150 DEG OR GREATER THAN 900 DEG C

X

E235

-

X

E236

AIRSPEED INPUT OPEN

X

E237

HEADING NOT VALID

X

E238

ENCODING ALTIMETER NOT VALID

X

E239

-

X

E240

-

X

E241

C/A RESET BUTTON PRESSED

X

E242

-

X

E243

-

X

E244

ERASE EEPROM SWITCH ON DURING TEST

X

E245

READ EEPROM SWITCH ON DURING TEST

X

E246

X

E247

X

E248

T Section 4 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-38.4

FAA APPROVED

NOTE

45.3 and 45.4, for performance information.

Due to safety margin designed into collective coupling system, Flight Manual performance may not be obtained during three cue coupled mode.

If Increased Maximum Continuous Power Kit (412-706-029)is not installed, refer to Flight Manual, BHT-412-FM-4, for performance information.

With Increased Maximum Continuous Power Kit (412-706-029)installed, refer to Flight Manual Supplement, Increased Maximum Continuous Power Kit (412-706029) with PT6T-3D Engine, BHT-412-FMS-

Refer to BHT-FMS- 412-53.4 for PT6T-3DE (30 minute OEI rating) engine performance. Refer to BHT-FMS- 56.3/56.4 for PT6T-3DF (30 minute OEI rating) engine performance.

17/18

BHT-412-FMS-38.4

T N I Section5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O No change from basic manual.

19/20

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

1-1. INTRODUCTION

Dual digital AFCS controls and indicators and related electrical systems, flight control system, and AFCS are partially of the described in this section data. For further manufacturer's information, including complete details and specifications, refer to the Honeywell SPZ-7600Integrated Flight Control System Pilot's Manual for the Bell 412.

1-2. DUAL DIGITALAFCS CONTROLSAND INDICATORS

All caution and advisory lights are dimmed using DIM switch on caution panel. Figures 1-1 through 1-11 depict the controls and indicators for the 4 axis AFCS/EFISinstallation.

1-3. ELECTRICALSYSTEMS

1-3-A. DC ELECTRICALSYSTEM

Refer to tables 1-1 through 1-3 for loading of nonessential, essential, and emergency buses.

1-3-B. AC ELECTRICALSYSTEM

I Refer to table 1-5 for bus loading.

1-4. FLIGHT CONTROL SYSTEM

BHT-412-FMS-38.4

The force trim system is activated by the master FORCE TRIM switch on the pedestal. Cyclic force trim can be released momentarily by pressing the FORCE TRIM switch on cyclic grip. Collective and antitorque force trim can be released by placing the CLTV/YAW TRIM switch (located on collective control head) in the OFF position or momentarily released individually by toggling in desired direction or jointly released by pressing straight down.

1-5. AUTOMATICFLIGHT CONTROLSYSTEM(AFCS)

The DFZ-7600 dual digital AFCS consists of two independent autopilot systems, AP1 and AP2, either of which is capable of four axis (pitch, roll, yaw, and collective) control, and flight director coupling. The system also incorporates other workload reducing features such as auto trim, heading hold, automatic turn coordination, and auto level.

Either autopilot can be operated in SAS (Stability Augmentation System) mode or ATT (Attitude Retention) mode. SAS mode provides short term stabilization for hands-on flying. ATT mode can be used for hands-off flight when coupled to flight director. The flight director provides steering commands which can be manually flown by the pilot or can be coupled to autopilots for fully automatic flight path control.

1-4-A. FORCETRIMSYSTEM 21

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

1-6.

BHT-412-FMS-38.4

EHSI

NAV 2 is primary navigation source for pilot EHSI and NAV 1 primary for copilot EHSI. Either side may select the other NAV unit as a secondary navigation source. During normal operations the primary navigation source should be displayed. Annunciation of selected navigation source is displayed in upper right corner of EHSI in white. The annunciator will illuminate amber when secondary NAV source selected or both EHSls display same NAV source. NAV 2 VOR course deviation and bearing pointer reference is derived from the pilot EHSI and may be displayed on either pilot or copilot EHSI.

NAV 1 VOR course deviation and bearing pointer reference is derived from the copilot EHSI and may be displayed on either pilot or copilot EHSI. NOTE

Loss of cross side heading or a heading miscompare, annunciated by an amber HDG in the lower left of EADI, may cause erroneous VOR course deviation and bearing pointer display when secondary navigation source is selected. Pilot shall disregard VOR course deviation and bearing pointer information or select primary NAV source.



21A/21B


BHT-412-FMS-38.4

r

IDLE STOP

START

Figure 1-1. Collective control panels (Typical)

22


Figure 1-2.

BHT-412-FMS-38.4

Overhead console, Typical (Sheet 1 of 2)

23

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W 0 O E I F V NLY O MANUFACTURER'S DATA

BHT-412-FMS-38.4

BASIC HEUCOPTER CONFIGURATION

i

----

301

MS1

TS

AC ____

. 051

F.

YRO

RAD

_

EFT5 AFC5

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ALT

14

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CPLT CPLT

XP

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ED P

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PILOT PILOT ATOT

9f]~ PP~s 0 REL FLOATS

CUT

PI.T

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AHRS EQUIPPED HEUCOPTERS

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_

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r-ACi WX 5

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RO

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LH

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AIR

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AFCS DATA SO

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A

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AIN

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HSI

AMD

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ALT1

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IRUS

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I

DC R

AC

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EFTS AFCS AlTTD .

SSNA--

NOT

I

C

A

NAVP COMM

ICS

CABN CS

SCDC

AHRS

RAO

LMR ALT2

FAN

S2

VO

NAVI VLF

VHF2

ENC IDENT'

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PIOT

PITOT

PR PR

CUT

NEL

AMRSCLOCK

PlOT

FOATS RELAY FORCEACS

PILOT STEP T

NOS2

412-FSMD-38.4-1-2-2

Figure 1-2. Overhead console, Typical (Sheet 2 of 2)

24

Rev. 2

16 AUG 0

BHT-412-FMS-38.4

MANUFACTURER'S DATA


AL-300 air data display (Typical)

412-FSMD38.4-1-4

Figure 1-4. MS-700 mode selector (Typical)

25

MANUFACTURER'S DATA

BHT-412-FMS-38.4


EADI (Typical)

EHSI

412.FSMD38.4-1-6

Figure 1-6.

EHSI (Typical)


BHT-412-FMS-38.4

412-FSMD38.4-1-7

Figure 1-7.

AP1

SAS/ATT

Radar altimeter (Typical)

AP2

TEST

FOSEL

COLL

CPL

RADALT VELHLD

412-FSMD38.4-1-8

Figure 1-8.

PC-700 autopilot controller

(Typical)

27


BHT-412-FMS-38.4

Figure 1-9.

Pilot/copilot

Figure 1-10. Pilot/copilot

display controller

(Typical)

EFIS aux switch panel (Typical)

BHT-412-FMS-38.4

MANUFACTURER'S DATA


Auxiliary navigation

panel (Typical)

29

BHT-412-FMS-38.4


MANUFACTURER'S DATA

DC electrical

system bus loading (Nonessential

buses)

NONESSENTIAL DC BUS 1


NAV-COMM SYSTEMS LF ADF (receiver) DME NAV 1 WX RADAR

NAV-COMM SYSTEMS VLF NAV VHF 2 COMM IDENT XPDR ENC ALT

LIGHTING SYSTEMS Position lights Console lights Pedestal lights Secondary instrument lights Copilot instrument lights Copilot map lights Landing light power Landing light control

LIGHTING SYSTEMS Anticollision light Baggage compartment light Cabin dome light

MISCELLANEOUS SYSTEMS Cabin heater Windshield heat controls (LH/RH) Windshield heat power (LH/RH)

MISCELLANEOUS SYSTEMS Air blower Vent blower

NOTES: 1.

Automatically sheds with one generator inoperative, but can be reenergized by placing the NON-ESNTL BUS switch in the MANUAL position.

2.

Automatically sheds for 30 and 90 minute flight with electrical power supplied by the battery only.


Table 1-2.

BHT-412-FMS-38.4

DC electrical system bus loading (essential buses)

ESSENTIAL DC BUS 1

ESSENTIAL DC BUS 2

NAV-COMM SYSTEMS Copilot ICS ICS Cabin RAD ALT 1

NAV-COMM SYSTEMS NAV 2 RAD ALT 2

FLIGHT INSTRUMENT SYSTEM EFIS NO. 1 BLWR EFIS NO. 1 SG/DC EFIS NO. 1 HSI EFIS NO. 1 ADI

FLIGHT INSTRUMENT SYSTEM EFIS NO 2 BLWR EFIS NO 2 SG/DC EFIS NO 2 HSI EFIS NO 2 ADI

(*** AHRS equipped helicopters ONLY ***): COPILOT AHRS

LIGHTING SYSTEMS Pilot instrument lights Engine instrument lights Utility lights

ENGINE 1 SYSTEMS Fuel transfer Fuel control heater Fuel boost Fuel crossfeed Fuel control Ignition Starter relay Particle separator Generator 1 reset

ENGINE 2 SYSTEMS Fuel control heater Fuel boost Fuel crossfeed Fuel control Ignition Starter relay Particle separator Generator 2 reset

NO 1 HYDRAULIC SYSTEM Hydraulic system power Hydraulic temperature indicator


CAUTION/WARNING SYSTEMS Engine 1 RPM

CAUTION/WARNING SYSTEMS Engine 2 RPM Rotor RPM

MISCELLANEOUS SYSTEMS Copilot pitot heater Copilot windshield wiper Governor control Nonessential bus power Hourmeter

MISCELLANEOUS SYSTEMS Pilot windshield wiper

FLIGHT CONTROLS AFCS NO 1 ADD/CAUT NOTES: 1.

Sheds for 90 minute flight with electrical power supplied by the battery only and with emergency load switch in EMERG LOAD position.

Rev. 2

16AUG00

31


MANUFACTURER'S DATA

Table 1-3. DC electrical system bus loading (emergency buses)

EMERGENCY DC BUS 1

EMERGENCY DC BUS 2

NAV-COMM SYSTEMS VHF COMM 1

NAV-COMM SYSTEMS Pilot ICS

(S/N 36248 and Sub) FLIGHT INSTRUMENT SYSTEM Standby attitude indicator Pilot AHRS

FLIGHT INSTRUMENT SYSTEM Standby attitude indicator LIGHTING SYSTEMS Searchlight power Searchlight control Pilot map light

ENGINE 1 SYSTEMS Fuel valve Fuel quantity indicator Fuel interconnect ITT indicator Oil temperature indicator Engine 1 torque indicator

ENGINE 2 SYSTEMS Fuel transfer Fuel valve Fuel quantity indicator Fuel interconnect ITT indicator Oil temperature indicator Engine 2 torque indicator

CAUTION/WARNING SYSTEMS Master caution panel Engine 1 fire detector Main fire extinguisher Baggage compartment fire detector

CAUTION/WARNING SYSTEMS Caution fail light Engine 2 fire detector Reserve fire extinguisher

MISCELLANEOUS SYSTEMS Mast torque indicator Idle stop release Cargo hook release Combining gearbox oil temperature indicator Inverter 1 power

MISCELLANEOUS SYSTEMS Passenger step XMSN oil temperature indicator Pilot pitot heater Emergency floats Hoist cable cut Inverter 2 power (S/N 36248 and Sub) AHRS FAN RELAY FLIGHT CONTROLS Force trim AFCS NO 2 CAUT

32

Rev. 2

16 AUG 00

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I FO V NLY O MANIJFACTURER'S DATA

BHT-412-FMS-38.4

Table 1-5. AC electrical system bus loading NO. 1 INVERTER

115 VAC BUS 1 Bus 1 voltmeter

115 VAC BUS 2 Bus 2 voltmeter No. 2 directional gyro

26 VAC BUS 1 No. 1 hydraulic pressure indicator Transmission oil pressure indicator Engine 1 oil pressure indicator Engine 1 fuel pressure indicator LF ADF (bearing pointers)

A115 VAC EMERGENCY BUS No. 2 attitude gyro

(S/N 36248 and Sub) zA115 VAC EMERGENCY BUS AHRS FAN

(S/N 36248 and Sub) 115 VAC BUS 2 Bus 2 voltmeter 26 VAC BUS 2 No. 2 hydraulic pressure indicator CBOX oil pressure indicator Engine 2 oil pressure indicator Engine 2 fuel pressure indicator No. 2 EFIS A 26 VAC EMERGENCY BUS No. 2 AFCS No. 2 rate gyro (S(N 36248 and Sub) a 26 VAC EMERGENCY BUS No. 2 AFCS ATTD 26 V2No.

NO. 2 INVERTER

115 VAC BUS 3 Weather radar stabilizer No. 1 directional gyro No. 1 attitude gyro

26 VAC BUS 3 No. 1 AFCS No. 1 rate gyro No. 1 EFIS

(S/N 36248 and Sub) 26 VAC BUS 3 No. 1 AFCS ATTD 26 V No. 1 ________________No. 1 EFIS

(S/N 36248 and Sub) 115 VAC BUS 3 Weather radar stabilizer

0_^^H~~~~~~~

NOTES:

A Does not shed when emergency load switch is in EMERG LOAD position.

Rev. 2

16AUG00

33

BHT-412-FMS-38.4

MANUFACTURER'SDATA

T N I Section2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O For handling and servicing Instructions related to components of dual digital Automatic Flight Control System (4-Axis)

with EFIS, refer to manuals supplied by manufacturers.

T N I R P N ROTORCRAFT O R I MANUAL FLIGHT T O F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4

SUPPLEMENT DUAL DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM, SEARCH AND RESCUE (SAR) WITH PT6T-3B/E AND PT6T-3D ENGINES 412-570-002

S/N 36020 -

36086

S/N 36087 AND SUB CERTIFIED 11 SEPTEMBER1992

This supplement shall be attached to Model 412/412EP Flight Manual when Dual Digital Automatic Flight Control System, Search and Rescue (SAR) modification has been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained In this supplement, or other applicable supplements, consult basic Flight Manual.

Bell Helicopter NOTICE COPYRIGHT COPYRIGHT 1998 INC. TEXTRON BELL HELICOPTER ANDBELLHELICOPTER TEXTRON INC.

8

ofTextron Inc A Subsidiary

TEXAS 76101 BOX412· FORTWORTH. POSTOFFICE

REVISION 3 -

26 MARCH 1998

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3& -39.4

NOTICEPAGE




PN

Rev.1

T N I R P N O R I O CT F U T O OD N R E P E E R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3

& -39.4

ROTORCRAFT FLIGHT MANUAL

SUPPLEMENT DUAL DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM, SEARCH AND RESCUE (SAR) WITH PT6T-3B/E OR PT6T-3D ENGINES 412-570-002

S/N 36020 - 36066 S/N 36067 AND SUB

TEMPORARY REVISION FOR STARTING PT6T-3D ENGINES AT TEMPERATURES OF 5°C (41°F) AND BELOW

TEMPORARY

16 OCTOBER 1994 REVISION - 19 MARCH 1997 REISSUE - 26 MARCH 1996

BHT-412.PMS-39.3 &-36.4

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Temporary ......................

Reissue ........................ Reissue .......................

LOGOF REVISIONS wMAR67

14Nov97 26MAR 98

LOGOF PAGES

Page Title........................ A-B ...................... 16.......................... 17..........................

Revision NO. Page Temporary Temporary Tempwary Temporary

TEMPORARY REVISION-II) MAR 97 REISSUE - 26MAR 98

Revision NO.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4

LOG OF REVISIONS

Original ........... Revision .......... Reissue...........0 Reissue ...........

0 .............. 1..............23 .............. 0..............18

11 SEP APR 13 JAN OCT

92 93 94 94

Revision .......... Revision .......... Revision ..........

............. 2 .............. 3.............

16 AUG 96 14 NOV 97 26 MAR 98

LOG OF PAGES

PAGE

Title ..................................... P N ...................................... A - B .................................. i- ii .................................... 1 ........................................ 2 - 7 ................................... 8 ........................................ 9- 15 .................................. 16- 17.................................3 18- 19.................................0 20 ....................................... 21 - 23 .................................

REVISION NO. 3 1 3 0 1 0 3 0 2 0

REVISION NO.

PAGE 24 25 34 35 36 37

....................................... - 33 ................................. ....................................... ....................................... ....................................... - 52 .................................

3 0 2 3 2 0

MANUFACTURER'S DATA

53 - 64 ................................. 65/66 ....................................

0 0

NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. Rev.3

A

BHT-41%FMS-39.3 8 -39.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOGOF APPROVEDREVISIONS

Temporary Reissue Reissue_.

19MAR97 _. 14NOY 97 98 _. _. 26MAR

/ CEKrlFlCATlON ROTORCRAFT OFFlCE FEDERAL Nvir,ONAOMlNlSTRAT,ON FT.WORTH, TX75193~0170 TEMPORARY RE”,SlON 97 s REISSUE - 2s MAR 98- 19MAR


LOGOF FAAAPPROVED REVISIONS

Original ........... 0 .............. Revision .......... 1 .............. Reissue...........0 .............. Reissue ........... 0 ............

11 SEP 92 23 APR 93 13 JAN 94 18 OCT 94

APPROVED:

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERALAVIATIONADMINISTRATION FT. WORTH,TX 76193-0170

B

Rev.3

Revision .......... Revision .......... Revision ..........

DATE:

1............. 2 .............. .............

16 AUG 96 14 NOV 97 26 MAR 98


FAA APPROVED

GENERAL INFORMATION

includes the The 412 SAR configuration Flight SPZ-7600 Integrated Honeywell Control System (DFZ-760 Digital AFCS, EDZ-705 EFIS, Primus 700 Weather Radar, DATA NAV III Information System, and AA300 Radio Altimeters), Global Wulfsberg GNS X Flight Management System, Racal D-91 Doppler, rescue hoist, and emergency to appropriate kits. Refer float manufacturer's publication for operations.

ABBREVIATIONS,ACRONYMS AND PLACARDING ADF

-

AGL AP APR BLWR CDI

-

CDU CLTV-YAW COMM CONT CPL CPLT CPU CRS DCPL DECEL DEG DH DME

-

DN DPLR DR EADI

-

EEPROM

-

EFIS

-

EHSI

-

Automatic Direction Finder Above Ground Level Autopilot Approach Blower Course Deviation Indicator Control Display Unit Collective-yaw Communication Control Couple Copilot Central Processor Unit Course Decouple Deceleration Degrees Decision Height Distance Measuring Equipment Down Doppler Dead Reckoning EFIS Attitude Director Indicator Electrically Erasable Programmable Read Only Memory Electronic Flight Instrument System EFIS Horizontal Situation Indicator

EMER ENC EXT FD FDSEL FMS

-

fpm FWD GNS

-

GPS

-

HDG HOV HSI

-

HTR IAS ICS

-

IDB IDENT ILS

-

LE LF LH LNAV MKR BCN MOT MSG NAV NDB NM PRI PWR RAD ALT REL RH RNAV SAR SBY ATT SG REV

-

STBY SYNC TACAN TAS TORQ/TRQ

-

Emergency Encoding Extinguisher Flight Director Flight Director Select Flight Management System Feet Per Minute Forward Global Navigation System Global Positioning System Heading Hover Horizontal Situation Indicator Heater Indicated Airspeed Intercommunication System Internal Data Base Identification Instrument Landing System Level Low Frequency Left Hand Long Range Navigation Marker Beacon Mark On Target Message Navigation Nondirectional Beacon Nautical Mile Primary Power Radar Altimeter Release Right Hand Radio Navigation Search and Rescue Standby Attitude Symbol Generator Reversionary Standby Synchronize Tactical Air Navigation True Airspeed Torque

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4 TRANS VEL HLD VLF VOR

-

VPU WP/WPT WSHLD WX XPDR XTRK

-

Transfer Velocity Hold Very Low Frequency VHF Omnidirectional Range VORTAC Positioning Unit Waypoint Windshield Weather Transponder Cross-track

FAA APPROVED


FAA APPROVED

1

Section

1-1.

1-5.

INTRODUCTION

CONFIGURATION

Limitations specified in Basic Flight Manual and supplements remain

1-5-A. REQUIREDEQUIPMENT

does not

Whenreferringto requiredequipmentlists

applicable.

Operation of SPZ-7600 system

relieve

flight

crew of

responsibility to monitor and comply with all helicopter limitations. For limitations of Emergency Floats, refer to BHT-412-FMS-2.

For limitations of Internal Hoist, refer to BHT-412-FMS-26.

in Basic Flight Manual note that EFIS replaces ADI/HSI; HP1/HP2.

and AP1/AP2

The EFIS composite

replace

mode is to be used in

IFR flight only after In-flight failure of an

1-3. 1-3.

TYPES TYP

OF OFESOPERATION OPERATION

EADI or EHSI. Use of composite

VFRflight is allowed.

mode for

IFR operations are not approved with:

1-5-B.

OPTIONAL EQUIPMENT

External cargo loads using SAR modes

NOTE

Hoist operations (unless using SAR

modes)

Any doors open (unless using SAR modes)

1-4.

FLIGHT CREW

The minimum flight crew for VFR or IFR operations consists of one pilot who shall operate the helicopter from the right crew seat. For IFR automatic approach and hover operations, two pilots are required. For overwater SAR operations, operator is required.

a hoist

Upon Initiation of a search and

rescue mission, the pilot in

command may determine

minimum equipment required to complete the mission using the automatic approach and hover

system. Two operational

EFIS systems

Either collective adjustable friction or collective force trim shall be operable for VFR flight. When operating over water: Hoist

Approved life vests for all occupants Life raft

Emergency floats Rev.1

1


1-5-D.

FAA APPROVED

SAR OPERATIONS (IFR)

1-11.

AMBIENT

TEMPERA-

TURES

NOTE

Offshore

approaches

under

For ground operations

IFR

when OAT is

between 48.9 and 51.7 °C (120 and 125 °F)

conditions shall be conducted In accordance with appropriate operating rules.

with doors and windows closed, and EFIS fans operating, EFIS operation is limited to 90 minutes.

In addition to basic equipment required for certification, the 412-705-006IFR kit, and items listed for SAR operations (VFR), the following items shall be operational for SAR operations (IFR):

1-23.

1-23-A.

FLIGHT OPERATIONS NORMAL VFR/IFR

Digital AFCS preflight test (level 1

Both autopilots

AP1 and AP2 shall be

engaged in ATT mode during flight.

Flight management system (Global GNS X) Collective force trim

minimum) shall be accomplished

prior to

first flight of the day or before planned flight into IMC. Flight director shall not be coupled to autopilots below 60 KIAS.

Flight director coupling to autopilot is not possible with only one autopilot engaged.

Doppler

Uncoupled ILS approaches shall not be

Two radio altimeters

flown in 3-cue DECEL mode.

Weather radar with ground mapping

1-23-B.

SAR MODES,VFR/IFR


Approach 2 operations shall be conducted into the wind.

Actual weight change shall be determined after kits are installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits. SAR operations shall be limited to OGE hover

SAR modes may be selected anytime within engagement limits.

weights.

descent shall be 60 KIAS.

1-7.

AIRSPEED

Minimum airspeed during Approach 1 and MOT prior to Approach 2 or final MOT

CAUTIO

Refer to figure 1-1 for airspeed limitations.

SAR MODES CAN BE COUPLED SINGLE AUTOPILOT; HOWEVER,

1-8. ALTITUDE Maximum

operating

SUBSEQUENT

pressure

altitude

is

20,000 feet - collective coupling limited to 16,000 feet pressure altitude.

2

ELECTRICAL

FAILURES OR AFCS FAILURES MAY RESULT IN TOTAL LOSS OF

AIRCRAFT STABILIZATION AND FLIGHT DIRECTOR GUIDANCE.


FAA APPROVED

Pilot shall have hands lightly on controls when coupled to SAR modes below 100 feet AGL dual autopilot or anytime when coupled single autopilot. NOTE

When collective is coupled, the FZ-706 Flight Control Computer limits engine ITT or MAST TORQUE based on the engine series or transmission configuration installed.

CAUTION

DURING NORMAL AFCS COUPLED OPERATION, SYSTEM MAY ALLOW TOLERANCES BASIC HELICOPTER LIMITATIONS TO BE EXCEEDED FOR SHORT IT REMAINS THE DURATIONS. RESPONSIBILITY OF THE PILOT TO MONITOR AND CORRECT FOR ANY EXCEEDANCES.

B/E ENGINES and BASIC TRANSMISSION. Engine power/transmission are limited to 810°C ITT or 100% MAST TORQUE when is below 60 KIAS (or in GO airspeed AROUND mode with airspeed below 105 Above 60 KIAS engine power/ KIAS). transmission are limited to 765°C ITT or 81% MAST TORQUE. B/E ENGINES and MAX CONTINUOUS POWER KIT. Engine power is limited to 810°C ITT when airspeed is below 60 KIAS (or in GO AROUND mode with airspeed below 105 KIAS), and 765°C ITT when MAST is above 60 KIAS. airspeed TORQUE is limited to 100% when airspeed is below 105 KIAS and 81% when airspeed is above 105 KIAS.

D ENGINES and BASIC TRANSMISSION. Engines are limited to 810°C ITT during all is Transmission modes of operation. limited to 100% MAST TORQUE when airspeed is below 60 KIAS (or in GO AROUND mode with airspeed below 105

KIAS). When airspeed is above 60 KIAS, transmission is limited to 81% MAST TORQUE.

D ENGINES and MAX CONTINUOUS POWER KIT. Engines are limited to 810°C all modes of operation. ITT during is limited to 81% MAST Transmission TORQUE when airspeed is above 105 KIAS and 100% when airspeed is below 105 KIAS.

CAUTION

THE FZ-706 FLIGHT CONTROL DOES NOT COMPUTER RECOGNIZE ENGINE FAILURE. FOLLOWING ENGINE FAILURE, COMPUTER WILL CONTINUE TO LIMIT REMAINING ENGINE TO THE APPROPRIATE LIMITS STATED ABOVE UNTIL PILOT COLLECTIVE DISENGAGES COUPLING OR MANUALLY COLLECTIVE OVERRIDES LIMITING. UNDER CERTAIN LOW THE AFCS OAT CONDITIONS COMPUTER MAY EXCEED SINGLE ENGINE TORQUE LIMIT.

1-24. FLIGHT MANAGEMENT SYSTEM(GLOBAL GNS X) CAUTION

THE PRESENTLY DEPLOYED GPS CONSTELLATION SATELLITE DOES NOT MEET THE COVERAGE AVAILABILITY AND INTEGRITY FOR CIVIL REQUIREMENTS NAVIGATION AIRCRAFT ARE USERS EQUIPMENT. CAUTIONED THAT SATELLITE AND ACCURACY AVAILABILITY ARE SUBJECT TO CHANGE.

Prior to flight, verify by CPU readout on initialization page that following approved 3


computer program (or subsequently approved program) has been installed and that applicable operator's manual is available: Program 52

The following are specific and individual limitations applicable to various sensor modes of GNS X. When all sensors are operating normally, composite navigation information is output. In this case, limitations which apply are a composite of the following list:

FAA APPROVED

reveals an error greater than 3 NM along-track or cross-track during RNAVoperation.

GNS X with only VPU sensor operable is approved for enroute and terminal RNAV operation within the National Airspace System. GNS X with only Loran-C sensor operable is approved for enroute and terminal RNAV operation in the National Airspace System under the following conditions:

GNS X with only VLF/Omega mode is approved for:

Within the area depicted in the referenced operators manual.

VFR/IFR enroute RNAV operation within the counterminous United States and Alaska in accordance with AC 20-130 or AC 20-101C.

During RNAV operation of the GNS X in the Loran-C only mode, additional navigation equipment required for the specific type of operation in the National Airspace System must be installed and operable.

During RNAV operation of GNS X, in VLF/Omega mode only, additional navigation equipment required for the specific type of operation in the National Airspace System must be installed and operable.

GNS X position information must be checked for accuracy (reasonableness) prior to use as a means of navigation and under the following conditions: Prior to each compulsory reporting point during IFR RNAV operation when not under radar surveillance or control.

At, or prior to, arrival at each enroute waypoint during IFR operation along approved RNAV routes. Prior to requesting off-airway routing and at hourly intervals thereafter during RNAV operation off approved RNAV routes.

GNS X position should be updated when a cross-check with other approved navigation equipment

GNS X position information must be checked for accuracy (reasonableness) prior to use as a means of navigation.

GNS X, with only GPS sensor operable, shall not be used for navigation, nor shall the composite position be used within a terminal area in the event GPS is the sole available sensor. The GPS sensor input is designed to enhance other sensor(s) input resulting in a composite position output. When the has sufficient GPS sensor measurements and acceptable sensor geometry (an accurate altitude may be required) it will contribute to the GNS X Under these position. composite the GNS X composite conditions, position is approved for enroute and terminal RNAV operation.

During periods of dead reckoning, navigation shall not be predicated on the use of GNS X as a means of operation in the National Airspace System. Following a period of dead reckoning, the helicopter position should be verified by


BHT-412-FMS-39.3 & -39.4

visually sighting ground reference points and/or using other navigation equipment such as VOR, DME, TACAN, Loran-C,

When latitude/longitude, transferred from IDB, is displayed on CDU, pilot will ensure that it is a reasonable position

Doppler, NDB, or radar fix.

for the requested identifier.

The GNS X is not approved for approach operation. When operating outside the magnetic variation model area (north of 70° north latitude or south of 60° south latitude), pilot must variation.

manually

insert

magnetic

Internal Data Base (IDB)

IDB must be updated to latest revision

every 28 days. Updating may be accomplished with a Global-Wulfsberg Systems update disk, or equivalent.

The fuel mode is for advisory purposes only, and it should not replace primary helicopter fuel quantity gages.

1-25. DATA NAV Refer to Global

Navigation

System

Operators Manual for data on Data Nav III GNS X equipment.

1-26.

WEATHER RADAR

Refer to Weather Radar Operator's Manual

for weather radar system limitations.

Update disks will be received by mail (to

subscribers) or obtained at authorized Global-Wulfsberg installation centers or update centers.

5


& -39.4

°

FAA APPROVED

NOTE: 1. Pedal critical wind azimuth-hovering with the relative wind within SEE these azimuth angles NOTE can result In inability to 1 maintain heading due to large left pedal requirements for certain wind velocities. 2. Longitudinal cyclic 90 ° critical wind azimuth 95 aft cyclic may be with limited longitudinal AFCS hardover.

SEENOTE

SEE NOTE

270

SEENOTE

180


IGECRITICALRELATIVEWIND AZIMUTH

14,000 FT. DENSITY ALTITUDE LIMITED FOR IGE MANEUVERS

412-FS39.3-1-1

Figure 1-1.

Maximum speed -

sideward and rearward flight, crosswind and tailwind at a hover

T N I Section 2 R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

2-1.

BHT-412-FMS-39.3 & -39.4

INTRODUCTION

Refer to Weather Radar Operator's

This section contains instructions and procedures for operating the helicopter from the planning stage, through actual

flight conditions, to securing the helicopter after landing.

applicable.

and are not applicable

to

normal, maximum, and

cautionary operation ranges for the helicopter and its subsystems are indicated placards.

2-2. FLIGHTPANNING

to ensure safe operations and to provide the pilot with the data to be used during flight. Essential weight, balance, and

performance information compiled as follows:

The instructions and procedures contained herein are written for the purpose of

The minimum,

procedures.

Each flight should be planned adequately

Normal and standard conditions are assumed in these procedures. Pertinent data in other sections is referenced when

standardization all situations.

by instrument

Manual

for weather radar system normal

markings and

Anytime an operating limitation

is

Check type of flight and destination.

should be

to be performed

Select appropriate performance charts

to be used.

2-2-A. TAKEOFF AND LANDING 2-2-A. TAKEOFF AND LANDING DATA

exceeded, an appropriate entry shall be made in the helicopter logbook. The entry shall state which limit was exceeded, the

Refer to Basic Flight Manual for takeoff and landing weight limits, and for takeoff and landing distance information.

attained, and any additional information

SAR operations are restricted to OGE

duration of time, the extreme value essential

in determining

action required.

the maintenance

The limits depicted on instrument markings and placards represent careful aerodynamic calculations that are substantiated by flight test data.

Refer to Section 1 of this and other

applicable supplements, and the Basic Flight Manual for subsystems restrictions.

hover weights.

2-2-B. WEIGHT AND BALANCE

Determine proper weight and balance of the helicopter as follows:

Consult

Manufacturers

instructions.

Data for

Compute takeoff and anticipated landing gross weight, check helicopter

center of 7


FAA APPROVED

gravity (CG) locations, and determine weight of fuel, oil, payload, etc.

All areas checked shall include a visual check for evidence of

Ensure loading limitations listed in Section 1 of Basic Flight Manual and

helicopter is flown near or over salt water or in areas of high

BHT-412-FMS-26 have not been

corrosion, particularly when industrialemissions.

exceeded.

2-3-A. BEFORE EXTERIOR CHECK

WARNING

Flight planning -

HOIST LOAD CAN CAUSE

LONGITUDINAL OR LATERAL CG GROSS TO EXCEED LIMITS.

Gross weight and CG -

GRAVITY SHALL BE COMPUTED

TO ASSURE LOADING WITHIN

Publications-Checked.

APPROVED LIMITS.

Portable fire extinguishers -

If additional loads are carried during hoisting operations, loads should be placed on side of helicopter opposite hoist position. operations

Battery -

are anticipated,

The pilot is responsible for determining

for

safe flight. Refer to figure 2-1 for preflight check sequence.


BOOST PUMPswitches -


Rev.3

OFF.

ON.

NOTE

If aft sumps fail to drain, the sump valves may be operated manually.

Forward and middle fuel sumps samples as follows:

Drain

Press-to-drain valves -Press. Fuel filters day as follows:

8

OFF.

Aft fuel sump drain buttons (left and right) - Press.

NOTE

The pilot walk-around and interior The pilot walk-around and interior following procedures. The preflight check is not intended to be a detailed mechanical check, 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.

Drain samples as

ENGINE 1 and ENGINE 2 FUEL switches - OFF.

PREFLIGHT CHECK

Is in condition

Connected.

Aft fuel sumps follows:

compute CG and gross weight with and without hoist load.

whether the helicopter

Condition I

and security.

NOTE

2-3.

Compute. Refer

to BHT-412-MD-3orBHT-412-FM-4.

CENTER OF AND CENTER WEIGHT AND WEIGHT OF

When hoist

Completed.

Drain before first flight of


FAA APPROVED

BOOSTPUMPswitches- ON.

Battery vent and drain tubes -

ENGINE 1 and ENGINE 2 FUEL switches - ON.

Searchlight and landing light - Stowed.

Fuel filter (left and right) -

Unobstructed.

Drain

samples. samples.

Antennas - Condition and security.

2. FUSELAGE -

ENGINE 1 and

OFF.

ENGINE 2 FUEL ENGINE 2 FUEL

CABIN LEFT SIDE.

Copilot door - Condition and operation;

glass clean.

OFF.



OFF.

ELT - Condition.

Position lights -

2-3-B. EXTERIORCHECK CAUTION

IF HELICOPTER

HAS BEEN

1. FUSELAGE--FRONT. -FRONT. Cabin nose -

wipers stowed.

Remote hydraulic filter bypass indicator Verify green.

unobstructed.

Cover(s) removed;

Static ports (left and right) Unobstructed.

Rotor blade (forward) - Remove tiedown. Visually check condition and cleanliness. Cabin nose ventilators Nose compartment -

Condition.

Passenger door - Condition and operation; glass clean. Condition of popout windows. wheels removed.

Emergency floats - Stowed; covers and supports clean and secured. NOTE

Ensure emergency floats have

Condition; all glass clean;

Pitot tube(s) -

of

Landing gear - Condition; handling

EXPOSED TO SNOW OR ICING CONDITIONS, SNOW AND ICE SHALL BE REMOVED PRIOR TO FLIGHT.

-

Check security

BOOST PUMP switches -

Unobstructed.

Condition; circuit

had periodic inflation and inspection.

Passenger step security.

Condition and

Underside

Emergency float activation switches Condition. Doppler housing - Condition.

3. FUSELAGE- AFT LEFT SIDE. Rotor blade (aft) -

Remove tiedown.

Visually check condition and cleanliness.

No. Check ment -- Check. No.11 engine compartment

breakers in; door secured. Radar dome -

Condition. 9


FAA APPROVED

No. 1 engine oil level - Verify actual presence of oil in sight gage. Visually check oil level and filler cap. Governor spring - Check condition. Engine fire extinguisher -

Check bottle

pressure gage and temperature range.

Combining gearbox filter - Check bypass indicator retracted.

GPS antenna and coaxial connector

Condition and security. Tail rotor blade -

Visually

cleanliness.

and

Tail rotor -

Avionics compartments - Security of components; check OVER TRQ warning

Tail skid - Condition and security.

overtorque.

-

Remove tiedown.

check condition

Oil cooler blower - Unobstructed.

flag (cat's eye) for indication of

Condition

and free

movement on flapping axis.

AFCS computers Secured; compartment door secured.

Intermediate (42 ) gearbox - Verify actual presence of oil in sight gage. Visually check oil level. Check filler cap, and chip detector plug for security.

Access doors and engine cowling -

Elevator - Condition and security.

Secured.

Tailboom -

Drain lines - Unobstructed. Engine exhaust ejectors removed; unobstructed. Oil coolers -

-

Covers

Unobstructed.

Tailboom secured.

Condition.

Baggage compartment

Aft compartment

Condition; access covers

Radio altimeter

antennas

-

directional alignment arrows. Tail rotor driveshaft covers -

Condition;

Secured.

CAUTION

DO NOT BEND ELEVATOR TRAILING EDGE TAB.

Elevator - Condition and security. Check for spring condition by moving elevator toward the leading edge down position.

-

Cargo

secured; smoke detector condition; door secured.

5. FUSELAGE -

4. TAILBOOM.

10

Tail rotor (90° ) gearbox - Verify actual presence of oil in sight gage. Visually check oil level, check filler cap and chip detector plug for security.

AFT RIGHT SIDE. -

Check unobstructed.

Tail rotor actuator - Check. Engine fire extinguisher -

Check bottle

pressure gage and temperature range.

Combining gearbox oil level - Verify actual presence of oil in sight gage. Visually check oil level.

Oil cooler blower - Unobstructed. No. 2 engine compartment -

Check.

No. 2 engine oil level - Verify actual presence of oil in sight gage. Visually check oil level. Filler cap secured. Access doors and engine cowling Secured.


BHT-412-FMS-39.3 & -39.4

Fuel filler secure cap.

Visually check quantity;

6. FUSELAGE -

CABIN RIGHT SIDE.

Passenger door -

Condition and

Verify actual

presence of oil in sight gage. Visually check oil level.

handling

Emergency floats - Stowed; covers and supports clean and secured.

Ensure emergency floats have inspection.

Passenger security.

step -

Secured.

Yokeassembly- Condition. Pitch horns -

Security and condition.

Elastomeric bearings, lead-lag dampers - Check general condition.

Security and

and

Condition

Droop restrainers condition.

-

Security and

Simple pendulum absorbers -pendulum Simple Security and condition.absorbers

Rotor blades - Visually check condition and cleanliness.

NOTE

had periodic inflation

Check general

Blade retention bolts proper latching.

Position lights - Condition.

Landing gear - Condition; wheels removed,

hub -

Mast retaining nut -

operation; glass clean. Condition of popout windows.

Transmission oil -

Main rotor condition:

Main driveshaft and coupling -

Condition and security where visible;

and

coupling for grease leakage and appearance of overheating.

Transmission oil filler cap - Secured.

Underside

No. 1 and No. 2 hydraulic

reservoirs -

Emergencyfloat activationswitches-

Visually check fluid levels; caps

Doppler housing - Condition.

Antenna(s) - Condition and security.

Condition.

Pilot door -

Condition and operation;

glass clean.

Check security


of

Combining gearbox oil filler cap -

Secured.

Anticollision security. security.

7. 7. CABIN CABIN TOP. TOP.

Hub and sleeve assembly condition.

Check

Swashplate, support assembly, and collective lever - Check condition. Main rotor pitch links condition.

secured.

Security and

light -

Condition and

No. 1 and No. 2 engine air intakesCovers removed, unobstructed; check

particle closed. doors closed. particle separator separator doors Engine and transmission transmission

cowling

cowling -

Secured.

Fresh air inlet screen -

Unobstructed.

Rotor brake reservoir cap -

Security. 11


BHT-412-FMS-39.3 & -39.4

2-4. INTERIOR AND

Hoist - Installed in desired position; check roof and floor stud adapters and locking collars properly secured.

PRESTART CHECK Emergency floats:

EMER INFLATION handle -

Down and

safetied. Nitrogen bottle - Secured and pressure within allowable limits for and ambient altitude pressure temperature

as shown on chart decal,

located on nitrogen bottle.

Emergency floats safety pin Removefor overwaterflight.

Cabin interior -

Cleanliness and

security of equipment.

Cargo and baggage (if applicable) Check security.

Protective breathing equipment (if installed) - Condition and properly serviced. For SAR operations - Check life vests, and life raft, installed and serviceable.

Boom actuator - Installed in proper position; all fittings secured. AIRCRAFT POSITION switch (hoist control box) - Set in proper position. Hook - Rotates freely on cable. Cable -

Check proper routing through

guide rollers, pulleys, and drums.

Gearbox oil levels -

Hoist operators pendant connectors secured.

ACTUATION

OF CABLE CUT

OF CABLE CUT

SWITCH ON HOIST CONTROL BOX CAN CUT CABLE, EVEN IF CABLE CUT CIRCUIT BREAKER IS

Opening or removing doors shifts

OUT.

helicopter center of gravity and reduces VNE. Refer to BHT-412MD-3 or BHT-412-FM-4.

removed.

CABLE

or

CUT switches

(pedestal

and

hoist)-- Off; covers safetied. Safety vests, tether straps, hoisting

slings, and litters -

2-4-A. HOIST INSTALLATION

Installed;

SWITCH ON PEDESTAL CAN CUT CAN CUT ON PEDESTAL

ACTUATION

NOTE

Secured

Check sight

glasses.

Condition and

secured or stowed.

CHECK

2-4-B.

HOIST OPERATION CHECK

HOIST PWR, CONT, and CABLE

If pilot plans to operate hoist, it shall be installed in forward right position.

circuit breakers -

BATTERY BUS 1 and BUS 2 switches ON. NON-ESNTL BUS switch -

12

CUT

In.

MANUAL.


BHT-412-FMS-39.3 & -39.4

ICS - Check communication pilot and hoist operator.

between

HOIST PWR switch - ON; check that green (power on) light and amber 20 FOOT CAUTION lights on hoist control box illuminate. Hoist OVERTEMP Press to test.

warning

lights

-

NOTE

HOIST (pilot and operator) and BOOM switches - Check all hoist functions for proper operation. Check that pilot HOIST switch overrides operator pendant HOIST switch. Hoist cable - Check for corrosion, kinks, flat spots, fraying, or broken strands.

Up limit switch actuator - Raise while hoist is reeling in and check hoist motor stops; then release and check hoist resumes operation. Hoist speed Reduce as cable approaches up limit. Check that hoist stops when hook reaches up limit without excess tension on cable. for flight;

HOIST PWR switch -

Rotor tiedowns -

Seat and pedals -

Removed and stowed. Adjust.

Seat belt and shoulder Fasten and adjust.

harness

-

Shoulder harness inertia reel and lock Check.

Maintain tension on hoist cable while reeling in and out to prevent slack.

Hoist stowed secured.

2-4-C. PRESTART CHECK

hook restraint

Flight controls - Friction collective lock removed.

NORMAL.

BATTERY BUS 1 and BUS 2 switches OFF.

EEPROM READ switch forward position. Lower pedestal circuit

Instruct ground crewmember to discharge helicopter static electricity before attaching load to hoist when possible.

-

Locked

breakers -

Collective control head switches CLTV/YAW TRIM switch -

NAVIGATION panel and DME OFF.

in

In.

OFF.

OFF.

CABLE CUT switches (pedestal hoist) - Off; covers safetied.

and

MKR BCN, DPLR,

COMPASS CONTROL switches (slave position).

-

MAG

As desired.

Radios and flight management system OFF. EFIS control panel -

As desired.

FUEL INTCON switch -

NOTE

as desired,-

Transmission chip detector indicators Check; reset if required.

ICS -

OFF.

NON-ESNTL BUS switch -

Directional control pedals - Check freedom of movement; position for engine start.


BOOST PUMP switches -

NORM. OFF.

OFF.

13

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4 FUEL XFEED switch -

FAA APPROVED

NORM.

Overhead circuit breakers -

ENGINE 1 and ENGINE 2 FUEL switches - OFF. PART SEP switches -

HYDR SYS NO. 1 and NO. 2 switches ON.

switch

Altimeters -

-

Spring

STATIC SOURCE switch -

FIRE EXT switch -

PRI.

OFF.

In (forward).

WIPERS switches -

OFF.

OFF.

CARGO RELEASE switch HEATER switch -

OFF.

OFF.

AFT OUTLET switch -

NON-ESNTL BUS switch loaded to NORMAL.

-

Spring

OFF.

OFF.

BATTERY BUS 1 and BUS 2 switches ON; check BATTERY caution light illuminates. NOTE

AFT DOME LIGHT rheostat and switch OFF. PITOT STATIC HEATERS switch -

NORMAL.

IF EXTERNAL POWER IS USED CONNECT (1000 AMPS MAXIMUM). CHECK 27 ± 1 VDC; ADJUST POWER SOURCE IF REQUIRED.

AND MAP LIGHT

FIRE PULL handles -

OFF.

ON.


Set.

APPROACH PLATE knobs - OFF.

ANTI COLL light -

ON, cover down.

Static check.

OFF.

POSITION LIGHT switch -


FORCE TRIM switch Instruments

-

OFF.

UTILITY LIGHT switch -

EMERG LOAD switch -

As desired.

ROTOR RPM audio loaded to AUDIO.

OFF.

EFIS MASTER switch -

NORM.

ENGINE 1 and ENGINE 2 GOV switches - AUTO.

STEP switch -

All LT rheostats -

In.

Test all lights when night flights are planned or anticipated. Accomplish light tests with external power connected or during engine runup.

ROTOR BRAKE lights - Test. Pull brake ON and check that both caution lights illuminate; return to OFF and check lights extinguish.

OFF.

NOTE


OFF.

STBY ATT switch - TEST; check standby attitude instrument light illuminates and OFF flag retracts momentarily, then switch OFF.

WSHLD HEAT switches -

OFF.

Rotor brake shall be off at all times when the engines are running.

FIRE 1 and 2 warning lights test switch - Press to test.


FAA APPROVED

BAGGAGE FIRE warning

light test

switch - Press to test (verify light flashes).

CYC CTR caution lights -

and follow through coupling may

be increased.

Press to test.

Throttles - Rotate engine 1 throttle full open, then back against idle stop.

Press.

engine 1 throttle to full closed, then

OVER TORQ caution light -

Check light illuminates and MAST TORQUEindicator reads 105+1%.

Actuate ENG 1 IDLE STOP release, roll

as desired. Repeat

apply friction

procedure using engine 2 throttle and ENG 2 IDLE STOP release. NOTE

CAUTION

When either IDLE STOP release is

activated, the appropriate idle stop plunger will not release if pressure is applied toward the

IF MAST TORQUE INDICATOR INDICATESAN ERROR GREATER THAN ± 1% FROM THE 105%

closed position

POSITION, THE MAST TORQUE

SYSTEM IS UNRELIABLE. ACTION IS MAINTENANCE

of the throttle.

Moderate frictions should be applied to

overcome follow-through

REQUIRED.

between throttles.

Clock - Set and running. Caution

PNL (All

panel TEST switch -

coupling

DECR for 8

RPM INCR/DECR switch seconds.

segments extinguish except CAUTION PANEL).

Caution panel TEST switch segments illuminate).

-

LT (All

Caution Panel RESET switch -

Press

Either engine may be started first; however, the following procedure

is provided for starting engine 1 first.

(MASTER CAUTION light extinguishes).

FUEL SYS test switch - FWD TANK, then MID TANK; note digital and needle indications. FUEL SYS DIGITS TEST switch PRESS (Digital display reads 888).


2-5.

ENGINE

ON.

START

-

2-5-A.

ENGINE 1 START

ENGINE 1 FUEL TRANS switch - ON; check NO. 1 FUEL TRANS caution light

extinguished.

ENGINE 1 BOOST PUMP switch -

ON;

check NO. 1 FUEL BOOST light extinguished. ENGINE 1 FUEL switch

-

ON. (FUEL

VALVE caution light will illuminate momentarily.)

If the helicopter has been cold soaked in ambient temperatures

ENGINE 1 FUEL PRESS indicator Check.

throttles will be difficult to move

Rotor - Clear.

of -18°C (0°F) or less, both

15

BHT-412.FMS-39.3 &-39.4 2-5-AENGINE 1START (CONT)

FAAAPPROV


-Alter GASPROD - Check 61f 1XRPM (N,) when throttle isatIdlestop. -AddAmblent temperalure 5-C(41’F)andbelow incorporate warmup procedurea asfou.xv*: * GOV switch - UANUAt. * ENG 1ihrotllo- Open slowly to maintain 77- 85%ENG RPM, a8 desired (mlnlmum of3minutes) -Alter Donatimxeaae ROTOR above 65%RPM untilXMSN oiltemperature ir above 15-C. -Add* ENG 1throttle - Reduce toidle . GOV switch -AUTO

TEMPORARY REVISION - 13MAR 97 REISSUE - 26MAR 98

26 Mar 1998


FAA APPROVED

CONDUCTED AT 97% ROTOR RPM OR ABOVE.

CAUTION

PROLONGED

EXPOSURE

TO

NOTE

AMBIENT TEMPERATURES OF 0°C (32°F) OR LESS MAY FREEZE MOISTURE IN THE ENGINE FUEL

On side slopes greater than five disregard CYC CTR degrees,

CONTROL SYSTEM. MONITOR

caution lights and position cyclic

ENG RPM (N2) DURING

as required.

COLD

WEATHER STARTING FOR OVERSPEED.IF AN OVERSPEED APPEARS IMMINENT, ABORT

Cyclic - Position as necessary to extinguish CYC CTR caution lights.

START AND CLOSE THROTTLE TO OFF POSITION.

NOTE

START switch - ENG 1 position. Observe starter limitations. Engine 1 ENGINE OIL pressure

-

Indicating.

Engine 1 throttle - Open to idle at 12% GAS PROD RPM (N1) minimum. Engine 1 ITT -

Monitor to avoid a hot


START switch RPM (N1). GAS PROD -

Check 61 ± 1% RPM (N1)

when throttle is at idle stop.

start. Maximum ITT during start is

1090°C, not to exceed two seconds above 960°C. If ITT continues to rise,

abort start by activating idle stop

NOTE

During extremely cold ambient idle rpm will be

release and rolling throttle fully closed.

temperatures,

Starter should remain engaged until ITT decreases. Do not attempt restart until

high and the ENGINE, XMSN, and GEAR BOX OIL pressures may

maintenance corrective accomplished,

has been

If engine fails to start, refer to FALSE START procedures this

section.


Ensure in full down

CLTV/YAW TRIM switch -

ON.

CAUTION

IF STICK CENTERING INDICATOR IS INOPERATIVE, SYSTEM GROUND OPERATION SHALL BE

Rev.3

exceed maximum limits for up to two minutes after starting. Warm

up shall be conducted at 77 - 85% ROTOR RPM at flat pitch.

NOTE

16

Off at 55% GAS PROD

Do not increase ROTOR above 85% RPM until XMSN OIL temperature is above 15°C.

ENGINE,

XMSN, and GEAR BOX OIL

pressures -

Check.

Engine 1 PART SEP OFF caution light Check extinguished.

CAUTION

DURING RPM INCREASE, ANY ABNORMAL INCREASE IN ONEPER-REV VIBRATION MAY INDICATE ONE OR MORE MAIN

FAAAPPROVED

BHT-4,2-FMC39.3 (L-39.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 2-5-B.ENGINE 2STAR1

- AfterGASPROD-Check 81e 1%RPM when engine 2throttle isonidlestop. -AddAmbient temperature 5’C(41-F)andbelow, incorporate worm upprocedures asfollowS: . GOV switch - MANUAL . EN02throttle - Open slowly lo maintain ‘l?- 95%ENG RPM, as desired (minimum of3minutes). I NOTE Duringextremelycold ambient temperatures, idleRPM willbehigh andengine oilpressures mayexceed maximum limitsforupto 2 mlnufes afterstarting.Warmup shallbe conducted at95%ROTOR RPM alflat pitch. . ENG 2thronle - Reduce toidle. . GOV switch - AUTO.

TEMPORARY REVlSlON 19MAR MAR 97 REISSUE -- 26 98

17

26 Mar 1998


FAA APPROVED

Engine 2 throttle - Open to idle at 12% GAS PROD RPM (N1) minimum.

DROOP RESTRAINERS TO DISENGAGE FROM

ROTOR FAILED

STATIC POSITION. VERIFY PROPER OPERATION PRIOR TO FLIGHT.

Engine 1 throttle - Increase to 77 - 85% ENG RPM (N2). Friction as desired.

Observe ITT

Engine 2 ITT limitations.

Monitor.

START switch RPM (N1).

Off at 55% GAS PROD

GAS PROD - Check 61 ± 1% RPM when engine 2 throttle is on idle stop.

NOTE

For ground operaion, maintain

Rotor RPM within allowable range.

CAUTION

Higher minimum Rotor RPM reduces blade flapping.

Rotor RPM-

Maintain 77 - 85%.

IF EXTERNAL POWER IS USED, PROCEED TO ENGINE 2 START. IF BATTERY WAS USED, PROCEED AS FOLLOWS:

ENGINE INDICATES 10 TO 15% HIGHER ENG RPM (N2) THAN THE ENGAGED ENGINE AND NEAR

ZERO

IF

TORQUE.

A

NONENGAGEMENT OCCURS, CLOSE THROTTLE OF THE

GEN 1 switch -ON. AMPS 1 -

ENSURE SECOND ENGINE ENGAGES AS THROTTLE IS INCREASED. A NONENGAGED

Check at or below 150 amps.

NONENGAGED ENGINE. WHEN THE NONENGAGED ENGINE HAS

STOPPED, SHUT DOWN THE

2-5-B. ENGINE 2 START

ENGAGED ENGINE.

ENGINE 2 FUEL TRANS switch

-

ON;


ON;

check NO. 2 FUEL TRANS caution light extinguished.

check NO. 2 FUEL BOOST light out (FUEL XFEED caution light will illuminate momentarily). ENGINE 2 FUEL switch

-

ON (FUEL

ENGINE 2 FUEL PRESS -

Check.

VALVE caution light will illuminate momentarily).

2 ENGINE

Indicating.

OIL

pressure

ENGAGEMENT

(HARD)

OCCURS.

DOWN BOTH MAINTENANCE

SHUT

ENGINES. ACTION IS

REQUIRED.

Engine 2 throttle - Increase slowly to match Engine 1 RPM (N2). Monitor tachometer and torquemeter to verify engagement of second engine.

Engine 2 ENGINEOIL pressure - Check.

ENG 2 PART SEP OFF caution Check extinguished.

light -

IF EXTERNAL POWER WAS USED DISCONNECT. GEN 1 SWITCH - ON.

- ENG 2 position. START switch Observe starter limitations. Engine

IF A SUDDEN

-

GEN 2 switch -

ON. (BATTERY BUS 1

will switch OFF automatically.)

Rev.3

17


FAA APPROVED

NOTE

2-5-E.

Only one BATTERY BUS switch (1 or 2) should remain ON with both

generatorsoperating.

Check all extinguished

Caution lights (except AFCS). ENGINE,

AMPS 1 and 2 -

Within

Within limits.

NOTE

an engine whenever it is deemed necessary to remove internally trapped fuel and vapor:

for

inadvertent

2-5-D. ENGINE FAILS TO START

When the engine falls to light off within 15 seconds after the throttle has been opened

to idle, the following recommended:

FUEL switch -

action is

IDLE STOP release - Actuate. Fully closed.

IGN circuit breaker -

Pull out.

Engage for 15 seconds, then

disengage.

UEL switch

-

OFF.

IGN circuit breaker-

Push in.

Allow the required cooling period for the starter before proceeding. Follow normal start sequence as described on preceding pages.

2-6. SYSTEMS CHECK 2-6-A. STICK CENTERING INDICATOR CHECK

INDICATOR CHECK

Starter - Disengage. FUEL switch -

ON.

BOOST PUMP switch - OFF.

ON as required.

ELT - Check transmission.

BOOST PUMP switch - ON.

Starter -

AMPS 2 will indicate a higher load than AMPS 1 until battery is fully charged.

Throttle -

procedure is used to clear

The following

Throttle - Fully closed.

XMSN, and GEAR BOX OIL

temperatures and pressures limits.

Radios -

DRY MOTORING RUN

CAUTION

OFF.

BOOST PUMPswitch - OFF.

DURING

After GAS PROD RPM (N1) has decreased to zero, allow 30 seconds for fuel to drain from engine. Conduct a

CYCLIC MOVEMENTS UNTIL XMSN OIL TEMPERATURE REACHES15°C.

DRY MOTORING RUN before attempting another start.

EXTREME

DO NOT DISPLACE CYCLIC MORE THAN 1.5 INCHES FROM CENTER TO CHECK THE SYSTEM. IF CYC CTR CAUTION LIGHTS DO NOT WITHIN THE 1.5 ILLUMINATE

INCH DISPLACEMENT, 18

COLD

AMBIENT TEMPERATURES LIMIT

THE


FAA APPROVED

2-6-C.

SYSTEM IS INOPERATIVE.

BEYOND POINT AT WHICH CYC

C TR

Throttles -

LIGHT

C A UTION

HYDRAULIC

CHECK

CYCLIC

DO NOT DISPLACE

PRELIMINARY

Set to idle.

ILLUMINATES.

NOTE

*NOTE

CYC CTR caution between inhibited ROTOR RPM.

Uncommanded

lights

control

movement

or motoring with either hydraulic

are

system off may indicate hydraulic system malfunction.

95 and 105%

Cyclic - Displace approximately 1.25 inch (31.7 mm) forward, aft, left and right.

HYDR SYS NO. 1 switch ON.

OFF, then

and

HYDR SYS NO. 2 switch -

OFF, then

Check CYC CTR caution

each

time

extinguishes

2-6-B.

light illuminates

displaced

when

ON.

when centered.

FORCE TRIM CHECK

Flight controls collective friction

Ensure cyclic and

set to minimum.

Cyclic and pedals -

2-6-D. ENGINE FUEL CONTROL CHECK CHECK Throttles (both) -

Idle.

Move slightly each

direction to check force gradients.

Cyclic FORCE TRIM release switch Press; check trim releases with switch pressed, released.

reengages

with

switch

FORCETRIM switch - OFF; check trim disengages and FT OFF caution light illuminates. FORCETRIM switch - ON, cover down.


collective

OFF. Check

and pedal force trim

disengage.


Flight controls illuminate collective).

Move to check PRYC

(pitch,

roll,

DO NOT ALLOW GAS PROD TO DECREASE BELOW 50% RPM (N1).

NOTE

In the vicinity

of 8000 feet

pressure altitude, GAS PROD RPM (N1) may not change significantly

when MANUAL fuel control is selected.

GOV switch (ENGINE 1 or 2) - MANUAL; observe a change in GAS PROD RPM

ON.

OUT OF DETENT caution

CAUTION

lights

yaw, and

(N1) and GOV MANUAL caution light

illuminates. Open respective throttle

carefully to ensure GAS PRODRPM (N1) responds upward, then return to idle position. Return GOV switch to AUTO. Check for a return to original GAS PROD RPM (N1) and GOV MANUAL caution

light extinguishes. Check second governor in like manner.

19


BHT-412-FMS-39.3 & -39.4 Throttles (both) -

5% ROTOR RPM.

2-6-E.

Increase slowly above

GOVERNOR CHECK

the interconnect

valve has been opened

by Bus No. 1 power and valve is functioning properly). FUEL INTCON switch -

OVRD CLOSE.

Check FUEL INTCON caution illuminates then extinguishes.

light

FUEL XFEED/INTCONtest switch -

TEST BUS 2 and hold.

IF HELICOPTER IS SITTING ON ICE OR OTHER SLIPPERY OR

LOOSE SURFACE, ADVANCE TO SLOWLY THROTTLES PREVENT ROTATION OF HELICOPTER.


FUEL INTCON switch -

gine 1 throttle

-Fully

open. Check

ENG stabilized at 95 + 1% RPM (N2).

Engine 2 throttle -

Fully open. Check

No. 1 engine increases 2% ENG RPM (N2) and both engines stabilize at 97 1% ENG RPM (N2).

OFF.

Check engine 2 FUEL PRESSdecreases, then returns to normal. Return switch to ON. OPEN. Check

FUEL INTCON caution light illuminates

then extinguishes. (This Indicates that the interconnect valve has been opened by Bus No. 2 power and that the valve is

functioning properly).

FUEL INTCON switch - NORM. Check FUEL INTCON caution light illuminates

RPM INCR/DECR switch - Full INCR. Check ENG does not exceed 101.5%

then extinguishes. FUEL XFEED/INTCON test switch -

RPM (N2). Set at 100% ENG RPM (N2).

NORM.

OVRDCLOSE.

2-6-F. FUEL CROSSFEED AND

FUEL XFEED switch -

INTERCONNECT VALVE CHECK

ENGINE 1 BOOST PUMP switch - OFF. Check FUEL PRESS drops to zero on affected system. Return switch to ON.

FUEL XFEED/INTCON test switch TEST BUS 1 and hold.

NOTE

FUEL XFEED switch -

After turning either boost pump off, FUEL BOOST caution

light

should illuminate on failed side only.

ENGINE 1 BOOST PUMP switch - OFF. Check Engine 1 FUEL PRESS decreases, then returns to normal. (This indicates

that the crossfeed

valve has been

opened by Bus No. 1 power and check

valve is functioning properly). Return switch to ON. FUEL INTCON switch -

OPEN. Check

FUEL INTCON caution light illuminates then extinguishes. (This indicates that 20

Rev. 2

Repeat procedure for ENGINE 2 BOOST PUMP switch. NORM.

2-6-G. ELECTRICAL SYSTEMS

CHECK

DC VOLTS AC VOLTS -

heck 27 +1 vdc. Check 104 to 122 vac.

AMPS 1 and 2 -

Check within limits.

EMERG LOAD switch - EMERG LOAD. Check that the following items remain o


FAA APPROVED AP2

ROTOR -

COMM 1 radio

Cyclic -

Pilot map light

Set to 100% RPM.

Centered, friction removed.

HYDR SYS NO. 1 switch - OFF. Check NO. 1 HYDRAULIC caution light and

Pilot ICS

MASTER CAUTION light illuminate and system 1 pressure drops to zero.

Essential engine instruments Essential flight instruments

Cyclic - Check normal operation by EMERGLOAD switch- NORMAL. Check moving cyclic forward, aft, left, and right that all loads dropped regain power. that all loads dropped regain powerapproximately Center cyclic. cyclic. one inch. inch. Center approximately one INV 1 switch - OFF; check INVERTER1 caution light illuminates. Check No. 1 and No. 2 AC VOLTS for indication

inverter

2 has assumed

Return INV 1 switch to ON. INV 2 switch -

that

all ac loads.

OFF; check INVERTER 2

caution light illuminates.

Check No. 1

and No. 2 AC VOLTS for indication

Collective - Check for normal operation by increasing collective control 1 to 2 inches. Repeat 2 to 3 times as required.

Return to full downposition. Note an increase

in force

move pedal in each direction.

required

to

that

inverter 1 has assumed all ac loads.

HYDR SYS NO.2 switch -

Return INV 2 switch to ON.

hydraulic system 2 remains operational,

STBY ATT switch -

and system 1 remains off.

ON.


NO. 1 HYDRAULIC

GNS X

-

OFF. Check

ON.

extinguishes,

normal

and system

pressure.

ON. Check

caution

light

1 regains

Check NO. 2

HYDRAULIC caution light illuminates and system 2 pressure drops to zero.

2-6-H. HYDRAULICSYSTEM

Cyclic - Check normal operation by moving cyclic forward, aft, left, and right

CHECK

approximately

one inch. Center cyclic.

Collective -

Check for normal

operations by increasing collective

The HYDRAULIC SYSTEMS CHECK is to determine proper

control 1 to 2 inches. Repeat 2 to 3 times as required. Return to fully down

actuators for each flight control system. If abnormal forces, unequal forces, control binding or

Pedals - Displace slightly left and right. Note the pedals are now hydraulically

operation of the hydraulic motoring are encountered,

be an indication malfunctioning

actuator.

flight

it may

of

control

a

Check collective fully down, friction removed.

FORCE TRIM switch -

position. boosted.


ON. Check

light NO. 2 HYDRAULIC caution extinguishes, system 2 pressure returns

to normal, and hydraulic system 1 remains operational.

OFF. 21

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4 Cyclic and collective

FAA APPROVED

friction

desired.

-

Set as

AFCS system -

Test as follows:

Verify AP1 and AP2 are disengaged.

NOTE

Cyclic and collective force trim

Cyclic and collective frictions should be set to minimum for SAR mode operations.

switches heck ON. eve 11 AFCS TEST switch - ON for level test.

FORCETRIM switch -

ON.

Verify AL-300 Air Data Command Display reads LE 1.

CAUTION


BOTH HYDRAULIC SYSTEMS SHALL BE OPERATIONALPRIOR TO TAKEOFF.

Pilot

NOTE closely

shall

controls

monitor

during test, taking

care

not to restrict control movements.

To prevent nuisance failures,

NOTE

hands and feet must remain off

System 1 will normally operate 10

to 20°C cooler than system 2.

2-6-J.

AFCS CHECK


centering

of AFCS actuator

is necessary.

the actuators


Cyclic -

ON.

NOTE

Verification

controls.

Failure of

to center could

result in reduced control margins

(pitch and roll) and abnormal control positions (pitch, roll, and yaw).

approximately one-half inch.


Collective - Verify movement up then down approximately one inch. Verify movement aft right

approximately neutral.

Pedals -

WARNING

one inch, then return to

Pedals - Verify aft and forward movement (left pedal then right)

Cyclic -

Pilot and copilot EFIS attitude indicators - Erect and set intensity as necessary.

Verify movement aft right

approximately neutral.

one inch, then return to

Verify aft and forward

movement (left pedal then right) approximately one-half inch.

Upon completion of test, verify AP1


22

and AP2 lights extinguish reads END 1.

and AL-300


FAA APPROVED

If error codes are present, press SBY (or cyclic NAV STBY switch) once to view each code until END 1 is displayed (refer to MANUFACTURER'S DATA, Section 2 of this supplement for list of error codes). Press SBY to continue with level 2 test if desired, or press TEST switch to exit AFCS test.

2-6-K. CABIN HEATERCHECK

GAS PROD - Check 75% RPM (N1) minimum (both engines).

Thermostat knob -

Full COLD.

NOTE

affects operation Heater performance. Refer to Hover Ceiling and Rate of Climb charts for HEATER ON in Basic Flight Manual.

HEATERswitch - As desired. VENT BLOWERswitch -

2-7.

As desired.

BEFORE TAKEOFF

transmission, Engine, gearbox, hydraulic, and electrical instruments Within operating ranges. and warning Caution Extinguished.

DO NOT OPERATE ABOVE 21°C OAT.

-

WARNING

HEATER

MODERATEFRICTION SHALL BE TO OVERCOME APPLIED FOLLOW-THROUGH COUPLING BETWEENTHROTTLES.

CAUTION

HEATER SWITCH SHALL BE TURNED OFF WHEN HEATED AIRFLOW DOES NOT SHUT OFF AFTER THERMOSTATIS TURNED TO FULL COLD, HEATER AIR LINE

LIGHT ILLUMINATES, OR CABIN HTR CIRCUIT BREAKERTRIPS.

HEATER switch -

lights

ON.

Throttles - Fully open. Adjust frictions.

RPM switch - Minimum beep (DECRfor 4-5 seconds). RPM (trim) switch for 4-5 seconds).

Minimum trim (-2

ROTOR- Check 95% RPMor greater.


ON.

Thermostat setting observe heated airflow.

Increase and

DEFOG lever - ON; check airflow is diverted from pedestal outlets to windshield nozzles. Return lever to OFF. AFT OUTLET switch - ON; check airflow distributed equally between pedestal outlets and aft outlets. Return switch to OFF.

RPM (trim) switch - Adjust to obtain matching torque at 100% ROTORRPM. Flight instruments - Check operation and set. Select required EFIS display. When flying IFR check as follows: IVSI -

Needle near zero.

Altimeter - Set and check. EHSI -

Check.

23

BHT-412-FMS-39.3 & -39.4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O EADI -

Check.

Radio altimeter - Zero altitude; OFF flag retracted; DH set as desired.

POSITIONlights -

As required.

ANTI COLL light - Check ON.

PITOT STATIC HEATERS switch - ON. Check ammeter for load indication. Leave ON in visible moisture when temperature is below 4.4°C (40°F); turn OFF If not required. Radio(s) -

Check functioning.

Flight management system - Set as desired. Refer to Global GNS X Operator's Manual. Auxiliary navigation panel DPLR, DME -

On.

MKR BCN,

Doppler - After 30 seconds check operation by pressing HOV switch on EFIS (DC-811) display controller and verify doppler is valid.

FLOATS switch - ARMED for over water operations; check FLOATS ARMED caution light illuminated. COCKPIT VOICE RECORDER TEST switch - Press and hold for three seconds. Verify meter indicates GOOD. Passenger seat belts -

Fastened.

Safety vests and straps secured to helicopter. All doors -

On and

Secured.

2-7-A. POWER ASSURANCE CHECK

Power assurance check should be performed daily in accordance with the Basic Flight Manual.

2-7-B. PROLONGED GROUND OPERATION

NOTE

NOTE

Amber Doppler display and no velocity bars - use light' lateral cyclic inputs to gently rock helicopter. Velocity bars displayed and amber doppler extinguished, validates doppler.

For prolonged ground operation, AFCS shall not be operated in ATT mode.

DME VOL - OFF switch (rheostat) will control cockpit display volume and VPU of GNS X.

Cyclic control into the wind.

Centered or slightly

AFCS - On and select ATT or SAS mode as desired (ATT mode shall be used during IFR flight; SAS mode recommended for ground operation, hover, and takeoff). FDSEL 1 or 2 -As

desired.

FORCETRIM switch - ON in ATT mode; as desired in SAS mode. 24

STEP switch - As desired (STOW for over water or hoist operations).

Rev.3

CAUTION

MINIMUM

ROTOR

-

97% RPM

FOR GROUND OPERATIONWITH STICK CENTERING INDICATOR SYSTEMINOPERATIVE. NOTE

For ground operation, maintain Rotor RPMwithin allowable range. Higher minimum Rotor RPM reduces blade flapping.

ROTOR -

77 - 85% or above, as desired.

Cyclic - Position as necessary to extinguish CYC CTR caution lights.


FAA APPROVED

Check torque required to

Hover power -

NOTE

On side slopes greater than five

hover at four feet skid height.

degrees, maintain 100% ROTOR RPM. CYC CTR caution lights are

inhibited.

NOTE

Downwind takeoffs

are not

recommended since the published takeoff distance performance will

2-8. TAKEOFF

not be achieved.

takeoff, pitch attitude must be adjusted commensurate with

iDuring

power application to prevent

DURING LIFTOFF TO HOVER, ANY ABNORMAL ABNORMAL INCREASE INCREASE IN IN ONE ONE

entering the AVOID area of the diagram. Torque Height-Velocity

INDICATE INDICATE ONE ONE OR OR MORE MORE MAIN MAIN

hover power while accelerating to

MAY

PER REV VIBRATION

DROOP RESTRAINERS RESTRAINERS ROTOR DROOP ROTOR FAILED TO DISENGAGE

FROM

STATIC POSITION. VERIFY

PROPER OPERATION PRIOR TO FLIGHT.

shall not exceed 15% above IGE Takeoff Climbout Safety Speed (Basic Flight Manual. )

Apply forward cyclic to

Cyclic control -

accelerate smoothly.

Collective pitch -

When AFCS is in ATT mode, the cyclic FORCE TRIM release and CLTV YAW TRIM REL switches

should be pressed before liftoff

(to trim actuators to center

positions) and should be held until desired climbout attitude is attained.

AP1 and AP2 -

As required.

FD1 and FD2 mode selectors Area -

SBY.

AIRSPEED - Within minimum for IFR).

Flight Director selected.

Verify Flight Director

Flight Director - After reaching 60 KIAS select modes as desired. NOTE

is increased, it may

be necessary to rematch engine

torques prior to reaching hover.

RPM (trim) switch TORQUE,

Adjust to obtain

during NAV, VOR APR, ILS or BC modes while either ARM or CAP,

temporarily deselect the mode. Reset the HSI for the new course

and intercept heading, then rearm for automatic capture.

GAS PROD RPM

(N1), or ITT, as required, and 100% ROTOR RPM.

(60 KIAS

To perform any course change, other than at station passage,

NOTE

matching

limits


Clear.

As collective

Adjust as desired

after reaching VTOCS(45 KIAS).

NOTE

Rotation

of the course

knob

during NAV, VOR APR, ILS or BC modes while either ARM or CAP,

outside the cone of confusion can 25


FAA APPROVED

falsely trigger the over station sensor resulting in degraded performance

FLOATS switch

-

OFF upon reaching

1000 feet above water or when over land; check FLOATS ARMED caution light extinguished.

retrimmed to new attitude. Attitude may also be adjusted with cyclic ATTD TRIM switch. For momentary attitude changes, manual

cyclic movement may be used; however, AFCS actuators may be saturated to limit authority when cyclic is moved manually.

CAUTION

CAUTION

WHEN OPERATING NEAR THE MAXIMUM MAST TORQUE LIMIT,

In three cue coupled climb (VS) or

MAY

Control Computer will allow

INADVERTENT

OVERTORQUE

OCCUR

DURING

MANEUVERING FLIGHT CONDITIONS INVOLVING TURNS AND/OR NOSE DOWN ATTITUDE

CHANGES. DECREASE POWER TO 90% MAST TORQUE PRIOR TO MANEUVERING HELICOPTER.

ENG -

97 to 100% RPM (N2).

altitude hold, the FZ-706 Flight

commanded airspeed to decrease to preserve commanded climb rate and altitude hold as a collective

limit is approached. NOTE

During single system autopilot

AIRSPEED- Within limits for flight altitude.

operation airspeeds

Engine, gearbox, and transmission

hold heading.

instruments-Withinlimits. NOTE

Refer to applicable operating rules for high altitude oxygen requirements.

2-9-A.

MANEUVERING WITH AFCS

(AP1 or AP2) at above 60 KIAS,

autopilot roll axis will normally

However, if

excxcessiveroll activity

is

experienced due to turbulence, pilot may wish to fly with yaw

force trim released. This will

cause autopilot roll axis to hold wings level. NOTE

Inflight

use of VG FAST ERECT

will disengage the switch respective

autopilot and decouple

Use normal pilot control techniques.

2-9-B. MANEUVERING WITH AFCS IN ATT MODE

Press cyclic FORCE TRIM release and CLTV YAW TRIM REL switches and maneuver as desired. Release switches

when desired attitude is reached.

Autopilot will hold attitude until

26

the automatic flight control modes.

NOTE

Under certain helicopter loading conditions, smoother flight may be obtained in turbulence by selecting 2-cue coupling instead

of 3-cuecoupling.


FAA APPROVED

2-9-C. FLIGHT MANAGEMENT SYSTEM (GNS X)

miles or more, updating is required in order to meet the enroute RNAV criteria.

Verify PROG 52 appears on bottom line of initialization page.

2-9-C-4. RATE-AIDING MODE

2-9-C-1. EFIS INTERFACE

Automatic True Airspeed (TAS) and heading are used for rate-aiding.

Course deviation indicator on EHSI provides center to full scale offset of five nautical miles crosstrack displacement.

If automatic TAS is not available and this event is combined with a condition where less that three VLF communication stations are being received, then rateaiding must be provided by a manual TAS entry. If automatic variation is not must be available, then rate-aiding supported by manual variation entry.

TO/FROMflag on EHSI is activated when a valid leg has been programmed. The legend changes from TO to FROM when passing the TO waypoint.

In the event that either TAS or variation entry is required, a message light will be displayed. The message light will remain on until action noted is taken.

When GNS X is selected as navigation source, FMS will be displayed in upper right corner of EHSI.

Master warning of system failure, dead reckoning, or no-valid leg is indicated by displaying a red X across the CDI on EHSI. Additionally, the MSG (message) light will flash on the CDU, and the remote MSG (message) light will illuminate whenever any of these conditions occur. When the MSG key is activated, a sensor message the indicating page will appear, occurrence.

2-9-C-2. AUTOPILOT COUPLING

GNS X may be coupled through autopilot when: FMS is selected as the navigation source on EFIS.

NAV is selected on FD mode selector.

2-9-C-3. ACCURACY CHECK

can be GNS X position information checked for accuracy (reasonableness) by reference to known ground position or VOR, DME, TACAN, Loran-C, NDB, or radar fix. When accuracy checks reveal the GNS X position to be in error by three nautical

2-9-D. SAR OPERATIONS

2-9-D-1. APPROACH TO HOVER (APR 1) NOTE

Primary pilot station for SAR modes will be right seat.

FORCE TRIM switch -

ON.

CLTC/YAWTRIM switch - ON. AP1 switch -

ON.

AP2 switch -

ON.

APR 1 switch illuminates.

Press. Verify light

Verify APR 1 annunciation on EADI and monitor 500 ft/min rate of descent. NOTE

RAD ALT mode will engage automatically, then automatically 27

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3& -39.4 disengage initiated.

when

FAA APPROVED

descent

is

either cyclic

*Selecting the MOT mode

NOTE

If selected from 200 ft to 2000 ft AGL, helicopter will decelerate to 60 KIAS and descend to 200 ft AGL.

Monitor automatic level off at 200 ft AGL according to radio altimeter, and airspeed hold at 60 KIAS. Confirm RAD ALT and IAS annunciation on EADI and APR 1 annunciation off. NOTE

During approach profile, airspeed may be captured at anytime by momentarily pressing cyclic FORCE TRIM release switch or longitudinal beep switch. RAD ALT may be captured anytime by momentarily pressing CLTV YAW TRIM REL switch or beeping CLTV-YAWTRIM switch UP or DN. During APR 1 engagement, heading can be changed by engaging HDG switch on mode selector and selecting desired heading with heading control, or by using left/right cyclic beep to change heading manually. NOTE

The APR1 mode will be cancelled by: · Completion procedure

of the

APR1

*Pressing the APR button on the SAR engage panel

· Selecting SBY through either MS700 Flight Director Mode Selector or remote decouple button on

2-9-D-2.

APPROACH TO HOVER (APR 2) NOTE

APR 2 can be engaged at altitudes between 20 ft AGL and 250 ft AGL, with airspeed less than 65 KIAS. APR 2 will not engage if these conditions are not met.

APR 2 switch illuminates.

Press. Verify light

HOV switch (EFIS display controller) On; monitor doppler ground velocity.

Note hover altitude displayed on AL-300. If desired, select new hover altitude between 50 and 150 ft AGL.

Monitor descent and check APR 2 and IAS annunciation on EADI. At 175 ft AGL, IAS hold will disengage and a slow deceleration to a hover will commence. NOTE

During approach, VEL HLD may be engaged anytime by momentarily pressing cyclic FORCE TRIM release switch or longitudinal beep switch.

During approach, roll axis will maintain a zero lateral doppler velocity. Reference can be changed using lateral beep switch. Heading can be changed using CLTV-YAWTRIMswitch in L or R position, and will produce a flat uncoordinated turn to selected heading.

Monitor capture of selected altitude on RAD ALT displays, primarily EHSI vertical display.


BHT-412-FMS-39.3 & -39.4

Monitor doppler groundspeed on EHSI hover display.

EADI - Ensure VEL HLD and RAD ALT annunciators are on and APR 2 is off

when approach 2 profile is complete. NOTE

The APR2 mode will be cancelled

by:

of the

*Completion

APR2

procedure

*Pressing the APR button on the SAR engage panel

*Selecting SBY through either MS700 Flight Director Mode Selector or remote standby button on either cyclic

doppler is valid.

To reduce size of MOT pattern,

MOT mode should be engaged at 200 to 500 feet AGL and 60 knots

groundspeed.

DPLR switch -

NORM.

MOT switch -

Press ON when directly

over target. Verify light illuminates.

Monitor descent and deceleration to 200

feet AGL and 60 knots groundspeed. NOTE

Hover waypoint and final

approach course into wind will be established at MOT mode engagement. Wind will continue to be updated during approach until TAS drops below 30 knots on final

approach,

*Selecting the MOT mode *Selecting

the climb

mode if

within the engagement parameters

2-9-D-3.

APR 2 -

indicated

display freezing

NO DOPPLER

If doppler signal is lost, continue with APR 2 profile. If doppler is acquired during approach, doppler will synchronize and

approach should continue normally. If

approach is completed without doppler, helicopter will obtain airspeed between 0 and 40 KIAS when VEL HLD is captured,

and lateral velocity up to 20 knots plus wind drift. Attempt to acquire doppler by beeping cyclic aft until airspeed goes to zero. If doppler is not acquired by this point, or visual contact is not made, initiate a go around.

by wind

speed and

direction and turning amber. Note any changes in wind direction and

speed prior to final approach. GNS X wind velocity will be displayed down to 5 knots, and EHSI wind vector and velocity will be displayed down to 1 knot.

When MOT is selected, helicopter

will turn shortest

direction

downwind, then capture 60 KIAS plus wind seen at start of approach. A pattern will be flown

so that helicopter arrives at final descen't point with 60 knot groundspeed at least 5 seconds prior to start of descent.

EHSI -

Adjust range as desired to keep

hover waypoint and final approach

2-9-D-4.

course in view. Mission equipment - Check.

MARK ON TARGET (MOT)

Final approach descent:

NOTE

MOT mode may be engaged anywhere

within

APR 1 limits

if

AL-300 -

Set desired hover height

between 50 and 150 feet AGL. 29


BHT-412-FMS-39.3 & -39.4

EFIS mode selector - HOV. Doppler groundspeed -

2-9-D-6.

HOVER

Monitor.

MOT light - Extinguished at completion of approach.

EADI annunciations

-

Check VEL HLD

OCCURS IF A HARDOVER DURING AUTOMATIC APPROACH

AND HOVER OPERATIONS, TAKE OVER IMMEDIATELY

NOTE

MANUALLY AND INITIATE CLIMB

The MOT mode will be cancelled by:

MODE. RETURN TO A SAFE ALTITUDE AND CORRECT MALFUNCTION PRIOR TO

*Pressing the MOT button on the

APPROACH

SAR engage panel

RESUMING AUTOMATIC OPERATIONS.

AND

HOVER

Selecting SBY through either MS-

700 Flight Director Mode Selector or remote decouple button on either cyclic *Selecting the VEL HLD mode if below 60 knots airspeed

*Selecting the RADALT mode

*Selecting the APR1, APR2, or climb mode if within specific engagement parameters for that

2-9-D-5.

MARK ON TARGET (MOT)

-

NO DOPPLER

IN VEL HLD MODE, WHEN SELECTION OF ANY FLIGHT DIRECTOR

MODES (PILOT OR COPILOT) WILL DISENGAGE VEL HLD.

DOPPLER IS REQUIRED FOR EXTENDED AUTOMATIC HOVER

OPERATIONS. IF DOPPLER GOES INTO MEMORY, AUTOPILOT WILL

MAINTAIN A CONSTANT

VELOCITY SEEN AT TIME OF DOPPLER IN-MEMORY. IF PILOT HAS NO OUTSIDE REFERENCE

Valid doppler is required to engage MOT mode. If doppler is lost, approach will continue unless doppler off time exceeds

WHEN IN ACCELEROMETER VELOCITYHOLD,THE CONDITION CANNOT BE MAINTAINED FOR

accuracy will be degraded when doppler Is

TO BE REACQUIRED.

approximately 8 seconds In a turn, or 32 seconds level. MOT final hover position

EXTENDED PERIODS OF TIME WAITING FOR DOPPLER SIGNAL

in memory. If doppler off time is exceeded, autopilot will disengage MOT mode and

TIME PRIOR TO ALLOWABLE INITIATION OF CLIMB MODE WILL

on final approach. To complete approach, engage APR 2 and refer to APR 2 - NO DOPPLER.

CONDITIONS, BUT A MAXIMUM OF 45 SECONDS IN A STABLE HOVER AND 30 SECONDS IF MOVING AT ANY VELOCITY IS IF PILOT IS RECOMMENDED.

engage IAS and RAD ALT after rolling out

VARY

DEPENDING

THE

ON

UNSURE OF HELICOPTER'S

ATTITUDE, AIRSPEED, OR DIRECTION, 30

CLIMB MODE CAN


FAA APPROVED

BE ENGAGED ANYWHERE IN AUTOMATIC HOVER ENVELOPE BY PRESSING COLLECTIVE GO AROUND SWITCH.

MONITOR POWER DURING OPERATIONS. HOVER COLLECTIVELIMIT IS 100% MAST TORQUE

OR

810°C

ITT.

MANUALLY ADJUST COLLECTIVE TRIM WITH COLLECTIVE RELEASED AS REQUIRED TO MAINTAIN POWER WITHIN TRANSITORY LIMITS. DO NOT PUSH AGAINST COLLECTIVE TRIM OR A SLOW COLLECTIVE OSCILLATIONMAY DEVELOP. NOTE

Coupled hover over water has been demonstrated up to Sea State 5.

Coupled hover is normally engaged automatically at completion of APR 2 or MOT profile, but can be engaged manually by selecting RAD ALT and VEL HLD modes on autopilot controller. Hover velocities can be commanded by using cyclic beeper in desired direction up to 10 knots laterally and aft, and 60 knots forward. The velocity reference will be displayed on AL-300 for five seconds for direction actuated by cyclic beeper switch. This will be indicated by a minus sign for aft or left, and no sign for right or forward. Cyclic beeper switch - Select direction and velocity desired. Press center of beeper switch to return hover reference to zero. NOTE

If in IMC, a maximum of 5 knots laterally and aft is recommended, depending on gross weight and wind conditions. If doppler return over water are conditions marginal, maintaining hover velocities less than 5 knots will reduce the possibility of doppler

memory excursions. Small hover corrections should be made by flying against trim, with zero velocity as the reference. Hover movements can be made by pressing cyclic momentarily FORCE TRIM release switch, but this is not recommended unless outside references are available. If cyclic FORCETRIM release switch is pressed, a new hover velocity reference will be established at the time of FORCE TRIM switch release.

CLTV-YAW TRIM UP and DN switch As desired. Select RAD ALT altitude as desired on AL-300. NOTE

Search and rescue operations will require extended hover in the Height-Velocity Diagram AVOID hover Automatic range. operations between 50 and 150 feet AGL require that pilot be attentive to controls in the event of a sudden engine failure. Altitudes below 50 feet AGL can be selected using CLTV-YAW TRIM UP and DN switch. Below 100 feet AGL, pilot is required to on have both hands lightly controls to respond to engine failure or autopilot malfunction.

Pilots HOIST EN switch - Press to transfer limited hover control to hoist operator (if desired). NOTE

Hoist operator has limited hover control by operating hover trim switch on hoist pendant in desired direction. White dot on HOIST EN switch will illuminate to indicate hoist operator is making control inputs. Helicopter can be flown up to 5 knots around a hover velocity reference established by the pilot. Pendant

31


FAA APPROVED

switch is held in a given direction until desired speed is reached using outside references, then released to capture that velocity. To return to zero velocity, switch is operated in opposite direction until velocity is zero, then released. Pilot must monitor doppler bars to ensure helicopter is responding properly. Pilot may take hover trim control away from hoist operator by pressing HOIST EN switch, or simply by overriding hoist operator's inputs using cyclic stick or cyclic beep switch.

2-9-D-7.

COLLECTIVE TRIM OR A SLOW COLLECTIVE OSCILLATION MAY DEVELOP. NOTE

When CLIMB mode is selected, helicopter will begin a slow, smooth transition into forward flight. Collective will maintain reference hover height during forward acceleration, then capture a 200 ft/min climb rate.

EADI - Check CLIMB annunciation.

CLIMB

NOTE NOTE

CLIMB mode is selected by pressing collective GO AROUND switch, or CLIMB switch located on instrument panel. CLIMB mode may be selected anywhere in SAR mode envelope if RAD ALT is between 20 and 200 feet AGL and less than 60 KIAS. Initiation of CLIMB mode will return helicopter to 200 feet RAD ALT and 60 KIAS hold. NOTE

CLIMB mode may be initiated at anytime during SAR mode hover when doppler returns are lost. CLIMB mode will function with or without doppler inputs to autopilot.

Collective GO AROUNDswitch - Press. CAUTION

MONITOR POWER DURING INITIAL FORWARD TRANSITION. COLLECTIVELIMIT IS 100% MAST TORQUE

OR

810°C

ITT.

MANUALLY ADJUST COLLECTIVE WITH COLLECTIVE TRIM RELEASED AS REQUIRED TO MAINTAIN POWER WITHIN LIMITS. DO NOT PUSH AGAINST

When helicopter reaches 60 KIAS, IAS hold mode will engage. Upon reaching 200 feet AGL, RAD ALT mode will engage. When both modes are engaged, CLIMB mode will extinguish.

During climb, VEL HLD may be engaged at anytime by momentarily pressing cyclic FORCE TRIM release switch or longitudinal beep switch. RAD ALT may be engaged at anytime by momentarily pressing CLTV YAW TRIM REL switch or beeping CLTV-YAWTRIM switch UP or DN. During climb, roll axis will maintain zero lateral doppler velocity. Heading can be changed using CLTV-YAWTRIM switch in L or R position, and will produce a flat uncoordinated turn to selected heading.

EADI - Monitor for completion of CLIMB mode.

2-9-D-8.

HOIST OPERATIONS

REFER TO APPROPRIATE FLIGHT MANUAL SUPPLEMENT FOR APPROVED LIMITATIONS AND PROCEDURES FOR HOIST OPERATIONS.


FAA APPROVED

HOIST switch -

NOTE

The Height-Velocity Diagram is not a limitation for internal hoist operations under an appropriate operating certificate.

HOIST PWR switch -

ON.

Pilots HOIST EN switch - Press to transfer limited hover control to hoist operator (if desired).

UP.

CAUTION

USE CARE TO PREVENT CABLE, HOOK, AND LOAD FROM FOULING ON FUSELAGE OR LANDING GEAR.

Maintain zero groundspeed clear of obstructions.

HOIST OPERATOR SHALL BE SECURED TO HELICOPTER WITH AN APPROVED SAFETY HARNESS DURING HOIST OPERATIONS.

Establish area.

hover over

Hoist hook restraint -

hoist

BOOM switch - IN to swing hoist boom and load into cabin, if possible. Take off into wind, if possible, allowing over adequate hoist load clearance obstacles if load is not internal.

operation

CAUTION

Removed.

BOOM switch (or pilot HOIST switch) OUT. NOTE

As hook nears the up or down limits, hoist speed automatically slows.

HOIST switch -

until load is

DOWN.

Discharge static electricity, when possible, and connect hook to load, observing allowable hoist load. NOTE

Pilot may need to provide assistance with roll cyclic when load is lifted off surface or released.

AIRSPEED WITH EXTERNAL LOAD IS LIMITED BY CONTROLLABILITY. CAUTION SHOULD BE EXERCISED WHEN CARRYING AN EXTERNAL LOAD, HANDLING AS CHARACTERISTICS MAY BE AFFECTED BY SIZE, WEIGHT, AND SHAPE OF LOAD.

AIRSPEED - As required for adequate not to exceed limits for controllability, (20 or 60 KIAS, as hoist operations applicable).

2-9-D-9.

LITTER HOISTING

of a transportation When emergency patient by litter is essential, every effort should be made to land helicopter for litter loading. Litter hoisting can be hazardous and should be accomplished only when a landing is not feasible and condition of patient precludes use of personnel hoisting sling.

33


FAA APPROVED

In addition to all other procedures contained herein, the following shall apply to litter hoisting operations. EMPTY LITTER

WARNING

HOISTING OR LOWERING AN EMPTY LITTER IN OPEN POSITION IS PROHIBITED. AND EMPTY LITTER SUSPENDED FROM HOIST IN OPEN POSITION CAN OSC I LLATE UNCONTROLLABLY IN ROTOR WASH AND CAN FLY UPWARD, STRIKING FUSELAGE OR TAIL ROTOR.

Prior to hoisting or lowering an empty litter, litter shall be closed and secured with straps. Litter should be suspended in a near-vertical position and sling straps should be drawn tight. ·

LOADED LITTER

CAUTION

A LOADED LITTER CAN ROTATE ABOUT CABLE DURING HOISTING. HOIST OPERATOR MAY HAVE TO GRASP LITTER SLING STRAPS TO CONTROL ROTATION AS LITTER APPROACHES LANDING GEAR.

2-10. DESCENTAND LANDING 2-10-A. DESCENT

WARNING

IF DIGITAL AFCS COMPUTER DOES NOT RECEIVE RADIO ALTITUDE SIGNAL WHEN IN A COUPLED ILS MODE, AUTOLEVEL WILL NOT OCCUR AND HELICOPTER WILL CONTINUE TO FOLLOW GLIDESLOPE SIGNAL.

- Adjust (minimum Flight controls friction for ATT mode operation).

LITTER PATIENT SHALL BE SECURED TO LITTER WITH SAFETY STRAPS.

HOIST HOOK CATCH SHALL BE SECURED WITH SAFETY PIN PRIOR TO HOISTING LITTER PATIENT. NOTE

If litter is suspended too far below hook, litter cannot be loaded into helicopter with hoist hook at upper limit.

Litter sling straps should be adjusted so that litter is 24 to 28 inches (61 to 71 centimeters) below hoist hook. 34

Rev. 2

AFCS desired.

Engage ATT or SAS mode as

FORCE TRIM switch - ON in ATT mode; as desired in SAS mode. Throttles ENG -

Fully open.

100% RPM (N2).

Flight path - Stay clear of AVOID area of Height-Velocity diagram in the Basic Flight Manual. For landing distance information in the event of engine failure during approach, refer to Basic Flight Manual. NOTE The FZ-706 Flight Control Computer does not recognize

T N I R P N O R I O T C .T F U O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4

FAA APPROVED

engine failure. Following engine

failure, the computer will continue to limit the remaining engine to

approximately normal cruise power until pilot disengages collective coupling or manually overrides collective limiting.

Irregular terrain may affect DECEL

capture

performance.

and auto-level

center cyclic and friction as necessary to extinguish CYC CTR caution lights. NOTE On side slopes greater than five degrees, disregard CYC CTR caution lights and position cyclic

as required.

FLOATS switch -

Weather radar or standby.

Both control panels off

ARMED when over

FLOATS switch -

water; check FLOATS caution light

Complete;

Stick centering check -

OFF; check

FLOATS

ARMED caution light extinguished.

2-11.

ENGINE SHUTDOWN

illuminated.

AP1 and AP2 - Disengage; autopilot lights extinguish.

STEP switch - As desired (STOW when operating over water).


Cyclic -

check

OFF.

Frictioned as desired. Maintain

cyclic stick as near center as possible at

CAUTION

all rotor speeds.

RUN ON LANDINGS MAY RESULT IN ROLL OSCILLATIONS WHILE

ON THE GROUND.

IF THIS

COLLECTIVE FULLY DOWN O AP1 AND AP2 DISENGAGING WILL STOP THE OSCILLATIONS.

NOTE

by

blade flapping

Minimize

maintaining highest rotor RPM (NR) within allowable range.

Reduce to 77 - 85% Rotor

Throttles -

RPM, as desired.

2-10-B.

LANDING

Collective pitch Pedals -

ITT -

Check for inadvertent

transmission.

Centered.

FORCE TRIM switch -

AFCS -

Stabilize for one minute.

Fully down.

ON.

SAS mode.

CAUTION

STBY ATTD switch EMERG LT switch -

OFF

DISARM.

Engine instruments -

Within limits.

IDLE STOP release

switch

-

ENG 1

position.

MINIMUM ROTOR - 97% RPM FOR GROUND OPERATION WITH

Engine 1 throttle -

SYSTEM INOPERATIVE.

decreasing.

STICK CENTERING INDICATOR

ITT

and

Fully closed. Check

GAS PROD

RPM

Rev. 3

(N1)

35

BHT-412-FMS-39.3 & -39.4



ON.

IDLE STOP release switch -

Pilot - Remain at flight controls until rotor has come to a complete stop.

ENG 2

position.

Lighting and miscellaneous switches OFF.

Engine 2 throttle -

Fully closed. Check

BATTERY BUS1 1 and BATTERY BUS and BUS BUS 2 2 switches switches -

ITT and GAS PROD RPM (N1) decreasing.

Collective down lock desired.


Secured as

OFF.

2-12. POSTFLIGHT CHECK


OFF.

ENGINE1 and 2 FUEL switches-

Emergencyfloats safety pin -

OFF.

ENGINE 1 and 2 BOOST PUMP switches - OFF. ENGINE 1 and 2 FUEL TRANS switches

-

If conditions following:

require,

Installed.

perform the

NOTE

Refer to Manufacturers Data Manual for additional information.

OFF.

Avionics-

OFF.

Check general condition of droop restraint system and verify that the droop restraint arms are engaged in the lower detent of the cam window.

Install main rotor blade tiedown socks on blades and secure to mooring points.

USE COLLECTIVE TO DO NOT USE SLOW ROTOR RPM. USE OF COLLECTIVE

TO SLOW ROTOR

CAN CAUSE EXCESSIVE FLAPPINGAND OR CONING. CAUTION

AVOID RAPID ENGAGEMENT OF ROTOR BRAKE IF HELICOPTER IS ON ICE OR OTHER SLIPPERY OR LOOSE SURFACE TO PREVENT ROTATION OF HELICOPTER.

Rotor brake -

As desired. Apply at or

below 40% ROTOR RPM. Return to stowed position after main rotor stops. 36

Rev.2

Install tall rotor tiedown strap and secure to vertical fin.

Install exhaust covers, engine inlet protective plugs, and pitot tube covers.


FAA APPROVED

Figure 2-1. Pref412.FS39.3-2-1

Figure 2-1.

Preflight check sequence

37


FAA APPROVED

Section 3

3-1. INTRODUCTION

OAT CONDITIONS,

For emergency procedures relating to emergency floats, refer to BHT-412-FMS-2; for emergency procedures relating to

ENGINETORQUELIMIT

hoist, refer to BHT-412-FMS-26.

3-3.

3-3-A.

AFCS

COMPUTER MAY EXCEED SINGLE ENGINE TORQUE LIMIT.

NOTE

SAR operations require flight

within the Height-velocity diagram

ENGINE

AVOID range (Basic Flight Manual.

SINGLE ENGINE FAILURE

Engine failure (OEI) while flying in these

conditions

requires

immediate pilot action to

maximize helicopter performance during recovery.

CAUTION

AIRSPEED above VTOCS(45 KIAS):

RUN ON LANDINGS MAY RESULT

RUN ON LANDINGSMAY RESULT ON THE GROUND. IF THIS

OCCURS,

LOWERING

COLLECTIVE

FULLY

DISENGAGING

DOWN OR

AP1 AND AP2

WILL STOP STOP THE OSCILLATIONS. OSCILLATIONS. WILL CAUTION

THE FZ-706

COMPUTER

FLIGHT

CONTROL

DOES

NOT

LIMIT REMAINING ENGINE TO THE APPROPRIATE LIMITS STATED IN SECTION 1 -

LIMITATIONS UNTIL PILOT DISENGAGES COLLECTIVE OVERRIDES LIMITING.

38

AIRSPEED -

Maintain 55 to 65 KIAS.

RPM switch - INC. Set ENG at 100% RPM (N2) if possible. Fly helicopter to safe enroute altitude.

RECOGNIZE ENGINE FAILURE. FOLLOWING ENGINE FAILURE, COMPUTER WILL CONTINUE TO

COUPLING

Collective - Reduce as required to maintain ROTOR RPM and power within

OR MANUALLY

COLLECTIVE

UNDER CERTAIN LOW

Refer to Basic Flight Manual for additional

AIRSPEED hover:

procedures.

below

VTOCS (45 KIAS), and

Collective - Reduce as required to maintain ROTOR RPM and power within OEI limits.

AIRSPEED -

Increase to VTOCS(45

KIAS), if possible.


BHT-412-FMS-39.3 & -39.4

If flyaway is not possible, or hover cannot be maintained, prepare for

accelerometers, or by other electrical malfunctions. Failure of No. 1 hydraulic

Refer to BHT-412-FMS-2

SAS or ATT mode functions. Failure of No.

ditching.

procedures.

ELECTRICAL

3-7.

3-7-A.

for additional

SYSTEM

AC POWER FAILURE

system will render AFCS yaw axis

inoperative but will not affect pitch or roll

2 hydraulic system will not affect AFCS.

If both autopilots are disengaged, the following procedures do not apply. 3-9-A.

AFCS FAILS TO ENGAGE

OR DISENGAGES

INDICATION:

NO. 1 or NO. 2 INVERTER caution light

illuminates. Possible loss of power to certain ac instruments (with no INVERTERcaution light).

INDICATION:

AFCS caution light illuminated. NO. 1 AUTOPILOT/NO. 2 AUTOPILOT caution light illuminated.

AP1 or AP2 off (light not illuminated).

PROCEDURE:

Possible erratic API indications

on AP1

illuminates, proceed as follows:

AC VOLTS - Check to determine that remaining inverter has assumed all ac loads. INV PWR circuit breakers -

Check in.

Possible ATT FAIL flag displayed on pilot or copilot EADI.

Possible illumination of NO. 1 or NO. 2 INVERTER caution light.

AP1 or AP2 switch (affected system) -

NOTE

Press to reengage autopilot.

If power is lost only to certain ac instruments, but INVERTER caution lights remain out, proceed as follows: AC FEEDERS circuit - Check in.

breakers (8 each)

During IFR flight, if both inverters fail, land

as soon as practical; or continue flight

If inverter 1 or 2 fails, AP1 or AP2

will disengage,

but can be

reengaged by pressing respective switch on AFCS control panel. PROCEDURE:

AIRSPEED less.


under VFR, if desired.

INV 1 and 2 switches - ON; check NO. 1 and NO. 2 INVERTER caution lights

3-9. AUTOMATIC FLIGHT

extinguished.

CONTROLS

Pilot and copilot

SYSTEM

The automatic flight control system can be affected by malfunctions of any of the vertical gyros, yaw rate gyros, inverters,

EADIs -

Check ATT

FAIL flags retracted, indicators functioning properly.

Check the following circuit breakers in: 39


FAA APPROVED

SAS/ATT switch -

CAUTION

illuminated.

Check ATT light

If AUTOTRIMcaution light illuminated -

DO NOT ATTEMPT TO RESET ANY CIRCUIT BREAKER MORE THAN ONCE.

Follow procedure FAILURE.

for

AUTOTRIM

If malfunction persists, follow procedure for AFCS FAILS TO ENGAGE OR

INV 1 PWR and INV 2 PWR

DISENGAGES.

AC FEEDERS (8)

NO. 1 and NO. 2 ESNTL BUS FEEDERS (ON MAIN DC)

3-9-C. AFCS HARDOVER OR ABNORMAL CONTROL DISTURBANCE

AFCS (No. 1 and No. 2)

NOTE


NOTE

Maximum demonstrated altitude

PILOT ATT SYS and CPLT ATT SYS

AP1 or AP2 switch (affected system) Press to reengage.

loss with an autopilot malfunction was 25 feet during a SAR coupled

hover or during ° a coupled ILS approach with a 3 glideslope.

If either autopilot will not reengage, or

PROCEDURE

if abnormal control disturbance occurs, proceed as follows:


WARNING

Disengage.

If IFR, land as soon as practical; or

continue flight under VFR, if desired.

IF AP1 OR AP2 FAILS

DISENGAGED,

OR IS

REDUCE

AIRSPEED TO 115 KIAS OR LESS.

If both autopilots fail to reengage,

Cyclic FORCETRIM release switch -

proceed as follows: AIRSPEED -

AIRSPEED As desired. As desired.

CLTV YAW TRIM REL switch -

If IFR, land as soon as practical;

continue flight under VFR, If desired.

or

Press.

Correct helicopter attitude with cyclic and cyclic FORCE FORCE then release release cyclic and pedals, pedals, then TRIM switch and CLTV YAW TRIM REL

switch.

3-9-B. AFCS FAILS TO HOLD ATTITUDE PROCEDURE:

40


Actuator position indicators -

FORCETRIM switch - Check ON. CLTV/YAW TRIM switch -

AIRSPEED less.

ON.

Check

both systems. If any API shows maximum

displacement

or erratic

operation of any actuator, switch affected autopilot OFF.


BHT-412-FMS-39.3 & -39.4

If IFR, land as soon as practical,

continue flight under VFR, if desired.

3-9-D.

or

AUTOTRIM FAILURE

If an attitude source fails, pitch scale, roll

pointer, and aircraft symbol will be painted blue and a red ATT FAIL will be displayed in the middle of the sphere.

3-14-A-2.

INDICATION:

AUTOTRIM caution light illumination. PROCEDURE:

GLIDESLOPE, EXPANDED LOCALIZER, AND RATEOF-TURN FAILURES


If any of these systems fail, the respective pointer will be removed and a red X will be drawn through the respective scale.

Attempt to reset AUTOTRIM light by cycling either AP1 or AP2 off, then ON.

3-14-A-3.

If fault condition cannot be cleared, use error code displayed on AL-300 to

Certain internal failures within the symbol generator will result in display being

to Refer axis. failed determine MANUFACTURER'S DATA, Section 2 of

INTERNAL FAILURES

replaced by a red X centered on EADI with notation SG FAIL.

this supplement for list of error codes.

3-14-B.


OFF.

3-14-B-1.

Continue flight with manual control of collective and pedals. ATT mode may still be used, including pitch/roll FD coupled operation.

For pitch or roll trim failures:

Pilot may choose to remain in ATT

mode, using cyclic FORCE TRIM release

switch and APIs to manually re-trim as required to keep actuators working near center; or, pilot may select SAS mode with FORCETRIM switch ON or OFF as desired.

EFIS MALFUNCTIONS

3-14-A. EADI FAILURES (RED ANNUNCIATIONS) 3-14-A-1.

ATTITUDE FAILURE

SAME ATTITUDE

SOURCE

Cycle either AP1 or AP2 off, then ON.

3-14.

EADI CAUTIONS (AMBER

ANNUNCIATIONS)

For yaw trim failures:

If pilot and copilot have selected the same attitude source, the attitude source will be annunciated in amber. 3-14-B-2.

COMMON SYMBOL GENERATOR

When in SG REV mode, one symbol generator is driving both pilot and copilot

display tubes. During this mode, a reversionary annunciation is given in amber indicating the display source. This annunciation appears in the upper left corner of the attitude sphere and will display SG1 or SG2, depending on source. 3-14-B-3.

DECISION HEIGHT

When radio altitude

is within

100 feet of

decision height, a white box will appear to the left of radio altitude display. When at or below decision height, an amber DH will

appear inside the white box.

41


BHT-412-FMS-39.3 & -39.4

3-14-B-4.

FLIGHT DIRECTOR

HDG FAIL is displayed at the top of the

FAILURE

compass card.

An amber FD FAIL warning will be displayed in the upper left quadrant of EADI. Also, flight

director

3-14-C-2.

FAILURE

command bars

and all mode annunciations will be removed.

COURSE DEVIATION OR GLIDESLOPE DEVIATION

A failure of these systems results in removal of course or glideslope pointer, and a red X being drawn through

NOTE

FD FAIL is also displayed while the Digital AFCS TEST mode is

scale.

The letter G will appear at the zero deviation position of vertical scale to identify invalid information.

engaged.

3-14-B-5.

Amber

DECISION HEIGHT FAILURE

dashes

will

replace

numerical

3-14-C-3.

INTERNAL FAILURES

Certain internal failures within the symbol generator will result in display being replaced by a red X centered on EHSI with

values of decision height display.

notation SG FAIL.

3-14-B-6. 3-14--6.

3-14-D. EHSI CAUTIONS (AMBER ANNUNCIATIONS)

RADIO ALTITUDE FAILURE

Amber dashes will replace numerical

values and rising runway will be removed from display. 3-14-C.

EHSI FAILURES (RED

ANNUNCIATIONS)

3-14-C-1.

HEADING FAILURES

3-14-D-1.

HEADING SOURCE AND NAVIGATION SOURCE

Whenever pilot and copilot share same heading source or select same NAV radio,

appropriate source annunciation will turn amber. Annunciation will also be amber if both sides are displaying cross-side shortrange NAV sources of the same type.

A failure in the heading system will result

WAYPOINT

in removal of bearing pointers and their

3-14-d-2

course deviation indicator and scale, and compass sync indicator. The digital course select and heading select readouts will be replaced with amber dashes and a red

An amber WPT annunciation will be displayed to the left of the NAV source annunciation to indicate the passage of a long range NAV waypoint.

annunciators, course select pointer,

42


FAA APPROVED

3-14-E. FAILURE DISPLAYS AND CORRECTIVE ACTION

The following table is provided as a quick

reference to certain failure displays and the recommended corrective action. ANNUNCIATION OR INDICATION

FAULT CONDITION

FLIGHT DIRECTOR

CORRECTIVE ACTION

Red X and SG FAIL or both displays blank.

Symbol generator failure.

Flight director inoperative (no FD cues).

Press SG REV switch to revert to cross-side display and flight

NAV modes may

No action required.

Display reverts to

Display controller

failure.

disengage.

Display goes blank, or color and picture quality failure.

Electronic display failure.


HDG or CRS pointers

Remote instrument

power-up

configuration

and

director.

Display Controller will not operate.

bearing pointers

revert to NAV 1 and NAV 2.

jitter.

Turn failed display DIM knob OFF to obtain Composite Mode on

remaining display.

controller failure.

None.

If only one side is

affected, SG REV may be used at pilot's

discretion.

Blue sphere and ATT FAIL on EADI.

Attitude gyro failure.

Command bars out of view.

Press ATT REV switch to revert to cross-side

Red HDG FAIL on EHSI. Loss of HDG, CRS, etc.

Directional gyro failure.

Roll command bar may bias out of view.

Press HDG REV switch to revert to cross-side source.

Red X over CDI scale; loss of

NAV invalid.

None.

Check NAV tuning and station range. Select

bearing pointer.

source.

alternate NAV source if required.

3-15. SAR MODE MALFUNCTIONS

APR1, APR2, MOT, VEL HLD, RAD ALT, or CLIMB. AP1 and AP2 shall be engaged in ATT mode.

All required equipment shall be operable prior to commencement of SAR modes 43


FAA APPROVED

During SAR coupled mode operation, failure or abnormal disengagement of either AP1 or AP2 will not decouple the system. AFCS operation will revert to the remaining autopilot without interruption.

Reengagement of the autopilot may be attempted without affecting mode operation.

NOTE

At airspeeds above 60 KIAS, command bars will provide cues for holding 70 KIAS and 700 fpm climb.

Loss of both AP1 and AP2 during APR 2 or VEL HLD/RAD ALT modes:

In the event of any succeeding failure of required equipment, the coupled SAR mode operation shall be terminated

pilot.

by the

Select CLIMB mode by pressing collective GO AROUND switch or instrument panel CLIMB switch.

Manually fly helicopter to follow pitch and roll command cue.

NOTE

The AFCS, if coupled

in APR 1

heading mode following

failure or

mode, will remain coupled to

abnormal disengagement of either AP1 or AP2.

NOTE

At airspeeds

PROCEDURE:

climb to 200 ft AGL.

Reengagement of AP1 or AP2 may be attempted after helicopter has

Attempt reengagement of affected system.

DUAL AFCS FAILURE

PROCEDURE:

conditions.

Failed autopilot - Engage.

3-15-C.

NOTE

During SAR coupled mode operation, failure or abnormal disengagement of both AP1 and AP2 will decouple the system.

Loss of both AP1 and AP2 during APR 1 mode:

RAD ALT FAILURE

If coupled to failed RAD ALT, collective

command cue will bias out of view. A red

X will appear over RAD ALT display, indicating collective axis is no longer coupled to AFCS. THE PILOT MUST ASSUME COLLECTIVE AND ALTITUDE CONTROL BY CROSS REFERENCING THE ANALOG DISPLAY.

Select CLIMB mode.

CONTINUATION OF AN UNCOUPLED IMC

Manually fly helicopter to follow pitch and roll command bars and collective

APPROACHIS NOTRECOMMENDED. PROCEDURE:

3-15-C-1.

44

below 60 KIAS,

command bars will provide cues for holding 60 KIAS and 200 fpm

3-15-A. SINGLE AFCS FAILURE

3-15-B.

bars and collective

SINGLE RAD ALT FAILURE


BHT-412-FMS-39.3 & -39.4

If cross-side RAD ALT is valid, collective may be re-coupled

by pressing

FDSEL

3-15-C-3.

WEATHER RADAR FAILURE

switch.

Refer to Weather Radar Operator's Manual

for weather radar emergency and malfunction procedures.

DUAL RAD ALT FAILURE

3-15-C-2.

Manually fly collective/altitude and discontinue SAR operations under IMC conditions.

Table 3-1. Warning lights (Red)

PANEL

WORDING

FAULT CONDITION

CORRECTIVE

RPM (with audio)

Rotor rpm at or below 95%

Adjust collective

INCR-DECR required.

or

RPM (without audio)

ACTION

pitch and/or

switch

RPM

as

or

Rotor rpm at or above 105%.

Refer to Engine Fuel Control Malfunction procedures.

45


FAA APPROVED

Table 3-2. Caution lights

PANEL WORDING

FAULT CONDITION

NO. 1 INVERTER/NO.

Failure

2 INVERTER

inverter;

of

ac

CORRECTIVE ACTION

power

Check INV PWR circuit breakers in. Reengage AP1 or AP2. During IFR flight, if both inverters fail, land as soon as practical; or continue flight under VFR, if desired.

or

NO. 2 INVERTER

EMERG

AUTOTRIM

Pitch,

-

Place EMERG LOAD switch in NORMAL position, if electrical load shedding is not required.

or yaw trim

Cycle either AP1 or AP2 off, then ON. If light does not extinguish, FAILURE AUTOTRIM follow procedures in this section.

autotrim

Turn CLTV/YAW TRIM switch ON. Check excessive adjustable collective friction. Cycle either AP1 or AP2 off, then ON. If light does not extinguish, continue flight with manual collective inputs.

LOAD

switch

EMERGLOAD position. roll,

inoperative.

Check ac voltmeters to determine inverter remaining that automatically assumed load for failed Inverter.

CLTV

Collective inoperative.

NO. 1 AUTOPILOT/ NO. 2 AUTOPILOT

Autopilot 1 or 2 inoperative. Engage affected autopilot. If unable to engage, continue flight on remaining system.

NO. 1 EFIS FAN/NO. EFIS blower 2 EFIS FAN inoperative.

46

fan 1 or 2 Pull EFIS NO. 1 or 2 circuit breakers if OAT is above 49°C (120°F). Push circuit breakers in prior to landing.


BHT-412-FMS-39.3 & -39.4

Table 3-3.

Advisory

lights

PANEL WORDING

FAULT CONDITION

CORRECTIVE

DCPL


Ensure that AP1 and AP2 are engaged in ATT mode. Engage flight director modes as desired. Ensure CPL switch and COLL switch are ON.

FT OFF


Check FORCE TRIM switch ON and FORCE TRIM circuit breaker in.

MSG

GNS X message.

Check GNS X CDU for message.

DR

No NAV sensor information for GNS X. System is in dead reckoning mode and is using only HDG and IAS to compute position.

Rely on other NAV sources.

BATT

GNS X is operating own battery.

Restore primary power if practical.

XTRK (Green)

Pilot selected XTRK course offset.

Informational.

WP (Green)

Helicopter seconds change.

Informational.

P:R:Y:C: DETENT

OUT

OF

on its

is within 30 of next leg

Pitch, roll, yaw, and/or collective out of detent.

ACTION

Release force trim of affected system and center control.

47


BHT-412-FMS-39.3 & -39.4

Table 3-4. Digital AFCS error codes

PWR UP x x x x x x x x x x x x x x x x x

PRE FLT

CONT TEST

AL-300 DISPLAY


BHT-412-FMS-39.3 & -39.4

Table 3-4.

PWR UP

PRE FLT

CONT TEST

Digital AFCS error codes

AL-300 DISPLAY

(Cont)


FAA APPROVED

Table 3-4.

PWR UP

PRE FLT

CONT TEST


(Cont)

AL-300 DISPLAY


X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X

E139 E140 E141 E142 E143 E144 E145 E146 E147 E148 E149 E150 E151 E152 E153 E154 E155 E156 E157 E201 E202 E203 E204 E205 E206 E207 E208 E209 E210 E211 E212 E213 E214 E215 E216 E217 E218 E219 E220 E221 E222 E223

X

E224

X X X

E225 E226 E227

YAW SERIES ACT MOVED COLL ACT MOVED PITCH TRIM (H) E/A ON PITCH TRIM (L) E/A ON ROLL TRIM (H) E/A ON ROLL TRIM (L) E/A ON YAW TRIM (H) E/A ON YAW TRIM (L) E/A ON TRIM ENABLE FAIL TRIM ONE-SHOT FAIL PITCH TRIM (H) E/A OFF PITCH TRIM (L) E/A OFF ROLL TRIM (H) E/A OFF ROLL TRIM (L) E/A OFF YAW TRIM (H) E/A OFF YAW TRIM (L) E/A OFF SERVO VALID FAIL TRIM END AROUND FAIL APON OTHER INPUT FAIL AL-300 KNOB MOTION AIRSPEED MORE THAN 60 KNOTS PRIMARY RADALT MORE THAN 20 FT ANALOG GROUND FAIL LATERAL ACCELEROMETER FAIL LONGITUDINAL ACCELEROMETER FAIL VERTICAL ACCELEROMETER FAIL ROLL ATTITUDE MORE THAN 10 DEG PITCH ATTITUDE MORE THAN 10 DEG YAW RATE GYRO MORE THAN 10 DEG/SEC DIRECTIONAL GYRO FAIL VERTICAL GYRO NOT VALID DIRECTIONAL GYRO NOT VALID PRIMARY RADIO ALTIMETER NOT VALID AIR DATA SENSOR NOT VALID SERIAL I/O NOT VALID OTHER AUTOPILOT NOT IN TEST MODE SELECTOR BUTTON(S) PRESSED AP CONTROLLER BUTTON(S) PRESSED SAR CONTROLLER BUTTON(S) PRESSED HARDWARE NOT VALID FOUR-WAY BEEP SWITCH PRESSED ENC ALT LESS THAN -600 FT OR GREATER THAN 12,000 FT BARO ALT LESS THAN -600 FT OR GREATER THAN 12,000 FT POD DURING TEST ROD DURING TEST YOD+YFTR DURING TEST

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Section V4 NLY O BHT-412-FMS-39.3 & -39.4

FAA APPROVED

Table 3-4. Digital AFCS error codes (Cont)

PWR UP

AL-300 DISPLAY


x x x x x x

E228 E229 E230 E231 E232 E233

x

E234

COD DURING TEST FTR DURING TEST CFTR DURING TEST TRIM SWITCH OFF DURING TEST COLL TRIM SWITCH OFF DURING TEST ENG TORQUE LESS THAN 5% OR GREATER THAN 30% ENG TEMP LESS THAN 150 DEG OR GREATER THAN 900 DEG C

x x x x x x x x x x x x x

E235 E236 E237 E238 E239 E240 E241 E242 E243

PRE FLT

CONT TEST

E244 E245 E246 E247 E248

x

NOTE

Due to the safety margin designed into the collective coupling

AIRSPEED INPUT OPEN HEADING NOT VALID ENCODING ALTIMETER NOT VALID DOPPLER ARINC 429 BUS NOT VALID LNAV ARINC 429 BUS NOT VALID C/A RESET BUTTON PRESSED

CYCLIC BEEP ACTIVE ON DURING TEST ERASE EEPROM SWITCH ON DURING TEST READ EEPROM SWITCH ON DURING TEST VELHLD SYNC SWITCH ON DURING TEST SECONDARY RADALT MORE THAN 20 FEET SECONDARY RADIO ALTIMETER NOT VALID

Manual Flight system, Performance may not be obtained during three cue coupled mode.

51/52

T N I Section R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3

MANUFACTURER'S DATA

& -39.4

No cchange from basic manual.

53

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-39.3 & -39.4

1-1. INSTRUMENTPANEL AND CONSOLES


54

MANUFACTURER'S DATA

Figures 1-1 through 1-5 provide instrument panel, collective control panel, pedestal, overhead console, and hourmeter panel.


MANUFACTURER'S DATA

BHT-412-FMS-39.3 & -39.4

Figure 1-1.

Instrument panel

55

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

BHT-412-FMS-39.3 & -39.4

PILOTSIDE

Figure 1-2.

56

Collective control panels


BHT-412-FMS-39.3 & -39.4

57


MANUFACTURER'S DATA

412-FSMD39.3-2-4-1

Figure 1-4.

Overhead console (Sheet 1 of 2)


BHT-412-FMS-39.3 & -39.4

412-FSMD39.3-2-4-2

Figure 1-4. Overhead console (Sheet 2 of 2)

59


MANUFACTURER'S DATA

412-FSMD39.3-2-5

Figure 1-5.

60

Hourmeter panel


BHT-412-FMS-39.3 & -39.4

1-2. ELECTRICALSYSTEMS 1-2-A. DC ELECTRICAL SYSTEM

Table 1-5. DC electrical system bus loading (Nonessential buses) NONESSENTIAL

DC BUS 1

A


NAV-COMM SYSTEMS LF ADF (receiver) DME NAV 1 WX RADAR

NAV-COMM SYSTEMS VHF 2 COMM VLF NAV IDENT XPDR ENC ALT LR NAV DATA NAV DPLR

LIGHTING SYSTEMS Position lights Console lights Pedestal lights Secondary instrument lights Copilot instrument lights

LIGHTING SYSTEMS Anticollision light Baggage compartment Cabin dome light

light

Copilot map lights

Landing light power Landing light control

MISCELLANEOUS SYSTEMS Cabin heater Windshield heat controls (LH/RH) Windshield heat power (LH/RH)

MISCELLANEOUS SYSTEMS Air blower

Vent blower Hoist power Hoist control

NOTES: 1. Automatically sheds with one generator inoperative, but can be reenergized by placing the NON-ESNTL BUS switch in the MANUAL position. 2. Automatically sheds for 30 and 90 minute flight with electrical power supplied by the battery only.

61


Table 1-6.

DC electrical

ESSENTIAL DC BUS 1

MANUFACTURER'S DATA

system bus loading (essential

buses)

ESSENTIAL DC BUS 2

NAV-COMM SYSTEMS Copilot ICS RAD ALT 1 ICS CABIN

NAV-COMM SYSTEMS NAV 2 RAD ALT 2

FLIGHT INSTRUMENT SYSTEM EFIS NO. 1 BLWR EFIS NO. 1 SG/DC EFIS NO. 1 HSI EFIS NO. 1 ADI

FLIGHT INSTRUMENT SYSTEM EFIS NO 2 BLWR EFIS NO 2 SG/DC EFIS NO 2 HSI EFIS NO 2 ADI LIGHTING SYSTEMS Pilot Instrument lights Engine instrument lights Utility lights

ENGINE 1 SYSTEMS Fuel transfer Fuel control heater Fuel boost Fuel crossfeed Fuel control Ignition Starter relay Particle separator Generator 1 reset

ENGINE 2 SYSTEMS Fuel control heater Fuel boost Fuel crossfeed Fuel control Ignition Starter relay Particle separator Generator 2 reset



CAUTION/WARNING SYSTEMS Engine 1 RPM

CAUTION/WARNING SYSTEMS Engine 2 RPM Rotor RPM

MISCELLANEOUS SYSTEMS Copilot pitot heater Copilot windshield wiper Governor control Nonessential bus power Hourmeter

MISCELLANEOUS SYSTEMS Pilot windshield wiper

FLIGHT CONTROLS AFCS NO 1 PWR AFCS NO 1 ADD/CAUT

NOTES: 1. Sheds for 90 minute flight with electrical power supplied by the battery only and with emergency load switch in EMERG LOAD position.


Table 1-7.

DC electrical

BHT-412-FMS-39.3 & -39.4

system bus loading (emergency

EMERGENCY DC BUS 1

EMERGENCY DC BUS 2

NAV-COMM SYSTEMS VHF 1 COMM

NAV COMM SYSTEMS Pilot ICS

buses)

FLIGHT INSTRUMENT SYSTEM Standby attitude indicator LIGHTING SYSTEMS Searchlight power Searchlight control Pilot map light

ENGINE 1 SYSTEMS Fuel valve Fuel quantity indicator Fuel interconnect ITT indicator Oil temperature indicator Engine 1 torque indicator

ENGINE 2 SYSTEMS Fuel transfer Fuel valve Fuel quantity indicator Fuel interconnect ITT indicator Oil temperature indicator Engine 2 torque indicator

CAUTION/WARNING SYSTEMS Master caution panel Engine 1 fire detector Main fire extinguisher Baggage compartment fire detector

CAUTION/WARNING SYSTEMS Caution fail light Engine 2 fire detector Reserve fire extinguisher

MISCELLANEOUS SYSTEMS Mast torque indicator Idle stop release Cargo hook release Combining gearbox oil temperature indicator Inverter 1 power

MISCELLANEOUS SYSTEMS Passenger step XMSN oil temperature indicator Pilot pitot heater Emergency floats Hoist cable cut Inverter 2 power FLIGHT CONTROLS Force trim AFCS NO 2 PWR AFCS NO 2 CAUT

63


Table 1-8.

MANUFACTURER'S DATA

AC electrical

system bus loading

NO. 1 INVERTER

115 VAC BUS 1 Bus 1 voltmeter

115 VAC BUS 2 Bus 2 voltmeter No. 2 directional gyro

26 VAC BUS 1 No. 1 hydraulic pressure indicator Transmission oil pressure indicator Engine 1 oil pressure indicator Engine 1 fuel pressure indicator LF ADF (bearing pointers)

26 VAC BUS 2 No. 2 hydraulic pressure indicator CBOX oil pressure indicator Engine 2 oil pressure indicator Engine 2 fuel pressure indicator No. 2 EFIS 26 VAC EMERGENCYBUS No. 2 AFCS No. 2 rate gyro

115 VAC EMERGENCYBUS No. 2 attitude gyro

NO. 2 INVERTER

115 VAC BUS 3 Weather radar stabilizer No. 1 directional gyro No. 1 attitude gyro

26 VAC BUS 3 No. 1 AFCS No. 1 rate gyro No. 1 EFIS

NOTES: 1. Does not shed when emergency load switch is in EMERGLOAD position.

1-3. FLIGHT CONTROL SYSTEM

1-4. AUTOMATICFLIGHT CONTROLSYSTEM(AFCS)

1-3-A. FORCE TRIM SYSTEM

The DFZ-760dual digital AFCS consists of two independent autopilot systems, AP1 and AP2, either of which Is capable of four axis (pitch, roll, yaw, and collective) control, and flight director coupling. The system also incorporates other workload reducing features such as auto trim, heading hold, automatic turn coordination, and auto level..

The force trim system is activated by the master FORCE TRIM switch on the pedestal. Cyclic force trim can be released momentarily by pressing the FORCETRIM switch on cyclic grip.

Collective and antitorque force trim can be released by placing the CLTV/YAW TRIM switch (located on collective control head) in the OFF position or momentarily released individually by toggling in desired direction or jointly released by pressing straight down.

Either autopilot can be operated in SAS (Stability Augmentation System) mode or ATT (Attitude Retention) mode. SAS mode for provides short term stabilization hands-on flying. ATT mode can be used for hands-off flight when coupled to flight director. The flight director provides steering commands which can be manually flown by the pilot or can be coupled to

64

T N I R P N Section O 2 R I O T F C U T O OD N R Section 3 E P E E R R C S N G O N Section 4 I R W O E I F V NLY O MANUFACTURER'S DATA

BHT-412-FMS-39.3 & -39.4

autopilots for fully automatic flight path control. With the addition of doppler navigation system, doppler sensor, and SAR engage controller, SPZ-7600provides automatic approach to hover, mark on

target, and climbout modes for search and rescue operations. Refer to Section 3 for a complete list of error codes.




65/66

T N I R P N O R I O T F ROTORCRAFT C U T FLIGHT O OD MANUAL N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-40

SUPPLEMENTFOR INCREASEDGENERATOR CAPACITY KIT (412-706-026)

CERTIFIED 29 OCTOBER1992

This supplement shall be attached to the Models 412 and 412EP Flight Manuals when the Increased Generator Capacity Kit is installed.

The informationcontainedhereinsupplementsthe informationof the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult the basic Flight Manual

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1994 BELL HELICOPTER TEXTRON INC AND BELLHEUCOPTER TEXTRON A DIVISION OFTEXTRON CANADA LTD ALLRIGHTSRESERVED

POST OFFICE

BOX 482 . FORT WORTH. TEXAS 76101

REISSUE -

5 OCTOBER 1994

T N I R P N O R 0 I O T F C U T 0 O OD N R E P E E R R C S N G O N I R W 0 O E I F V NLY 0 O BHT-412-FMS-40

Additional copies of this publication may be obtained by contacting: Commercial Publication Distribution Center Bell Helicopter Textron P.O. Box 482 Fort Worth. Texas76101-0482


LOG OF REVISIONS

Original .......... Reissue ..........

0 0

.......... ..........

29 Oct 92 05 Oct 94

LOG OF PAGES

REVISION NO.

PAGE

Title ......................... NP .......................... A - B ....................... i/ii .......................... 1-18 ......................

PAGE

REVISION NO.

0 0 0 0 0

NOTE

Revisedtext is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

A


FAA APPROVED


Original .......... Reissue ..........

0 0

........... ...........

29 Oct 92 05 Oct 94

APPROVED:

MANAGER



FAA APPROVED

INTRODUCTION

The Increased Generator Capacity Kit increases the amperage limit

from 150 to 200 amps on each generator. The incremental

performance losses for the additional 50 amps each is presented in

Section 4.

i/ii

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O WEIGHT/CG

LIMITATIONS

GENERATOR

LIMITATIONS

Continuous operation Maximum for operation above 15.000 feet density altitude Maximum

0 to 200 amps 150 amps 200 amps

DURING SINGLE GENERATOR OPERATION. ELECTRlCAL LOADS SHALL SE REDUCED BEFORE RESTORING POWER TO NONESSENTIAL BUS TO ENSURE GENERATOR LOAD LIMIT IS NOT EXCEEDED.

AMMETER

0 to 200 AMPS

Co”tin”o”s operation

150 AMPS

Maxilwm ‘0, operation above 15,000 fl Ho

200 AMPS

Maximum

INSTRUMENT MARKINGS

T N Section 3 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-40

FAA APPROVED

ELECTRICALPOWERFAILURES

DC POWER FAILURE

If nonessential bus power is required, proceed as follows: Switch off all unnecessary equipment.

INDICATIONS:

DC GENERATORcaution light illuminates.

All lighting and avionics on nonessential buses inoperative.

PROCEDURE:

GEN FIELDand GEN RESETcircuit breakers Check in.

CAUTION

DO NOT SET NON-ESNTL BUS SWITCH TO MANUAL BEFORE TURNING OFF UNNECESSARY EQUIPMENTTO ENSUREGENERATOR LOAD LIMIT IS NOT EXCEEDED.

NON-ESNTLBUS switch - MANUAL.

GEN switch (affected generator) then ON.

RESET,

If generator remains inoperative, proceed as follows:

DC AMPS - Monitor.

Equipment switches - As desired/off as necessary to maintain generator load below maximum limit.

GEN switch (affected generator) - OFF.

If No. 2 Generator failed:

BATTERYBUS 2 switch - OFF; BATTERYBUS 1 switch - ON.

2

NOTE

During single engine operation, avoid generator load above 150 amps to attain climb performance presented in basic Flight Manual.

FAA APPROVED

T Section 4 N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-40

PERFORMANCE VARIATIONS

Performance variation charts are provided to determine hover and climb performance decrements due to the additional power requirements for the generators when operating at 200 amps each.

The charts are organized into three performance sections according to helicopter configuration and respective flight manual and supplements to which the charts apply.

PERFORMANCE SECTION APPLICATION

PERFORMANCE SECTION

HELICOPTER SERIAL NUMBERS

FLIGHTMANUAL/ SUPPLEMENT

Section 4A

33001-33107

None

*BHT-412-FM-1

Section 4B

33108-33213 36001-36019

None

*BHT-412-FM-2

33001-33107

Increased Gross Weight and Takeoff Horsepower (412-075-008-111)

36020 - 36086 AND 36087 AND SUB

None

33108-33213 36001-36019

Improved Hover Performance Modification (412-570-001-103) or Increased Maximum Continuous Power Kit (412-706-029)

BHT-412-FMS-34.2

Increased Maximum Continuous Power Kit (412-706-029)

BHT-412-FMS-41

Section 4C

33001-33107

*

EQUIPMENT REQUIRED

BHT-412-FMS-19.1

*BHT412-FM-3

*BHT-412-FM-4

BHT-412-FMS-41

Basic Flight Manual or appropriate optional equipment supplement. 3

T N Section 4A I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-40

TWIN ENGINE HOVER AND RATE OF CLIMB DECREASEDUE TO 200 AMP GENERATORLOADS.

Enter appropriate chart with pressure altitude and OAT to determine whether or not performance reduction is required. If applicable, decrease performance data in basic flight manual or appropriate optional equipment supplement as indicated on chart (rate of climb reduction of 30 feet per minute or hover gross weight reduction of 50 pounds out of ground effect or 60 pounds in ground effect).

FAA APPROVED


BHT-412-FMS-40

HOVER PERFORMANCEVARIATION FOR BHT-412-FM-1

TAKEOFFPOWER GENERATOR200 AMPS (EACH)

14000

13000 12000

11000

10000

9000

Figure 4A-1.

Hover performance variation - takeoff power

5


BHT-412-FMS-40

HOVER PERFORMANCE VARIATION FOR BHT-412-FM-1

MAXIMUM CONTINUOUSPOWER GENERATOR200 AMPS (EACH)

14000 13000

Figure 4A-2.

Hover performance variation - maximum continuous power


BHT-412-FMS-40

CLIMB PERFORMANCEVARIATION FOR BHT-412-FM-1


14000

13000

Figure 4A-3.

Climb performance variation - takeoff power

7

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 14000 13000 12000 11000

10000

9000

Figure 4A-4.

8

Climb performance variation - maximum continuous power

FAA APPROVED

T N Section 4B I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-40

BHT-412-FM-2 BHT-412-FMS-19.1

TWIN ENGINE HOVER AND RATE OF CLIMB DECREASEDUE TO 200 AMP GENERATORLOADS.

Enter appropriate chart with pressure altitude and OAT to determine whether or not performance reduction is required. If applicable, decrease performance data in basic flight manual or appropriate optional equipment supplement as indicated on chart (rate of climb reduction of 30 feet per minute or hover gross weight reduction of 50 pounds out of ground effect or 60 pounds in ground effect).

9

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O HOVER PERFORMANCEVARIATION

FORBHT-412-FM-2AND BHT-412-FMS19.1 TAKEOFFPOWER GENERATOR200 AMPS (EACH)

14000

13000 12000 11000

9000

w LU

8000

Figure 4B- 1. Hover performance variation - takeoff power

10


BHT-412-FMS-40

HOVER PERFORMANCEVARIATION

FOR BHT-412-FM-2 AND BHT-412-FMS-19.1 MAXIMUM CONTINUOUSPOWER GENERATOR200 AMPS (EACH)

14000 13000

12000 11000 10000

-

0 -40

FORPOINTSBELOWCURVE NOPERFORMANCE LOSS. -30

Figure 4B-2.

-20

-10

0


11


BHT-412-FMS-40

14000

13000

12000

11000

9000

- 2200 - 2000

OAT - C

Figure 4B-3.



BHT-412-FMS-40

CLIMB PERFORMANCEVARIATION

FORBHT-412-FM-2AND BHT-412-FMS-19.1 MAXIMUM CONTINUOUSPOWER GENERATOR200 AMPS (EACH)

14000

13000 12000 11000

10000

Figure 4B-4. Climb performance variation - maximum continuous power

13


BHT-412-FMS-40

BHT-412-FM-3

TWIN ENGINE HOVER AND RATE OF CLIMB DECREASEDUE TO 200 AMP GENERATORLOADS. Enter appropriate chart with pressure altitude and OAT to determine whether or not performance reduction is required. If applicable, decrease performance data in basic flight manual or appropriate optional equipment supplement as indicated on chart Irate of climb reduction of 30 feet per minute or hover gross weight reduction of 50 pounds out of ground effect or 60 pounds in ground effect).

14


BHT-412-FMS-40


FOR BHT-412-FM-3, BHT-412-FM-4, BHT-412-FMS-34.2,

AND BHT-412-FMS-41

TAKEOFF POWER GENERATOR 200 AMPS (EACH)

14000 13000

- - -- -

12000

HEATEROFF HEATERON WINTERIZATION KIT

FOR POINTS ABOVE CURVE SUBTRACT 50 POUNDS FOR OGE HOVER AND SUBTRACT 60 POUNDS FOR IGE HOVER FROM BASIC FLIGHT MANUAL OR SUPPLEMENTDATA

11000

10000

9000

FOR POINTS BELOWCURVE NO PERFORMANCELOSS.

Figure 4C-1.

Hover performance variation - takeoff power

15


FAA APPROVED


FOR BHT-412-FM-3,

BHT-412-FM-4, BHT-412-FMS-34.2.

AND BHT-412-FMS-41


14000

13000

12000 11000

Figure 4C-2.



BHT-412-FMS-40



AND BHT-412-FMS-41


14000 13000

12000

11000 10000

FOR POINTS BELOWCURV

Figure 4C-3.


17


FAA APPROVED



AND BHT-412-FMS-41


14000

13000 12000 11000 10000

Figure 4C-4.

18

Climb performance variation - maximum continuous power

T N I R P N A O R I FLIGHT MANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-41.3

MODEL EL

ROTOR

~ROTORCRAFT

SUPPLEMENT

INCREASED MAXIMUM CONTINUOUS POWER KIT WITH PT6T-3BE ENGINES (412-706-029)

36020 THROUGH 36080 CERTIFIED 14 AUGUST 1992

This supplement shall be attached to the Model 412 Flight Manual (BHT-412-FM-3) when the 412-706-029 Increased Maximum Continuous Power Kit has been installed.

* Helicopters S/N 36081 through 36086 are equipped

with PT6T-3D engines.


Db^~~ _______Bell ,

NOTICE COPYRIGHT COPYRIGHT 2000 ® INC. BELL HELICOPTER TEXTRON INC. TEXTRON ANDBELLHELICOPTER

of Textron Inc. A Subsidiary

TEXAS 71101 BOX42 *FORTWORTH, POSTOFFICE

ADIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

Helicopteriii

|REVISION

2 -5

26 APRIL 1995

SEPTEMBER 2000


NOTICE PAGE


NP

SEP Q2o000

BHT-412-FMS-41.3

T N I * 7 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODEL

ROTOR CRAFT FLIGHT MANUAL

IN

SUPPLEMENT SED MAXIMUM CONTINUOUS K T WITH PT6T-3BE ENGINES (412-706-029) GH 36080

TEMPO tIO N MAST TORQUE SI.ACONDITIONER

ASB 41

This

Tempora

Revision

exndc

perormance and povies new per

e

compliance with ASB 412-7-90.

Insert these temporary revision pages opposite like pages in Flight Manual after compliance with ASB 412 DO NOT remove existing pages from Flight Manual

Information contained herein supplements information of basic Fight Manual. For Umitations, Procedures, and Performnce Data not contained in this supplement, consult basic Right Manual.


Bell Helicopter * i 1

COPmRIGT *TIC coptfIhi

ASubsidilry ot tlxtrln irn

2000

BO ELL HELICmPT TTII,

ADISIm . TEXTRMA. A

ARLWs sRar

C

ADI

26 APRIL 1995

TEMPORARY REVISION -8

AUGUST 1997

CANCELED - 5 SEPTEMBER 2000




LOG OF PAGES

Revision No.

Page

Revision No.

Page

Title ........................ Temporary A-

...................... Temporary

5-8 ....................... Temporary 9/10 ........................ Temporary

NOTICE

This Temporary Revision has been canceled. Subject data has been incorporated in BHT-412-FMS-41.3 at Revision 2. Remove and discard all Temporary pages dated 8 August 1997 from this supplement including these canceled pages.

APPROVED:

MANAGER

DATE:

/


A


SEP

52000

T N I R P N O 0D R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-41.3

LOG OF REVISIONS

Original...........0 ............. 14 AUG 92 Reissue...........0 .............. 26 APR 95

Revision .......... ............. 08 AUG 97 Revision .......... 2 ............. 05 SEP 00

LOG OF PAGES

REVISION NO.

PAGE

Title- .................................. 2 NP......................................0 AB .................................. 2 i/ii ......... ... ....................... 0

*^^b~~~

REVISION NO.

PAGE

1-4...................................0 5 -8 .................................. 2 9/10 ..................................... 2

~~NOTE


O520 SEP

SEP 05 200nnn

Rev.2

A



O Original ...... Reissue ...........

............... 14 AUG 92 .............. 26 APR 95

APPROVED:

DATE:

MANAGER

ROTORCRAF CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

B

Rev. 2

08 AUG 97 Revision .......... 1 ........ Revision .......... 2.............. 05 SEP 00

SEP 05 2000

SEP

5

2000


BHT-412-FMS-41.3

GENERAL INFORMATION

The Increased Maximum Continuous Power Kit (412-706-029) provides for Increased hover and climb performance when the new mast torque indicating system is Installed.

Weight limitations data are provided for helicopters of early configuration which have been modified appropriately.

i/ii

FAA APPROVED

BHT-412-FMS-41.3

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 1-6. WEIGHTAND CENTEROF GRAVITY

Actual weight change shall be determined after components are installed and ballast readjusted, If necessary, to return empty weight CG to within allowable limits.

1-14. TRANSMISSION

CAUTION

WHEN OPERATING NEAR THE MAXIMUM MAST TORQUE LIMIT, INADVERTANT OVERTORQUE MAY OCCUR DURING MANEUVERING FLIGHT CONDITIONS INVOLVING TURNS AND/OR NOSE DOWN ATTITUDE CHANGES. DECREASE POWER TO 90% MAST TORQUEPRIORTO MANEUVERINGHELICOPTER.


1-14-C.MASTTORQUE 1-14-C-1. TWIN ENGINE OPERATION Maximum with airspeed up to 105 KIAS

100%

Maximum with airspeed above 105 KIAS

81%

1

T N I R P N O R I O T F C U T O OD N R E P E E R R C Section2 N-S G O N I R W O E I F Y V L Section 3 N O BHT-412-FMS-41.3

FAA APPROVED

TORQU

TORQUE MAST

GREEN

Figure 1-1. Instrument

10to81%

markings




FAA APPROVED

BHT-412-FMS-41.3

T N 4 I Section R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 4-1. INTRODUCTION

The performance data presented herein are derived from the engine manufacturer's specification power for the engine less installation losses when the 412-706-029 Increased Maximum Continuous Power Kit is installed. These data are applicable to the basic helicopter without any optional equipment which would appreciably affect lift, drag, or power available.

4-5. HOVERCEILING

Adequate cyclic and directional control are available at the gross weights allowed by the Hover Ceiling IGE charts in relative winds up th 35 knots from any direction at or below 3000 feet HD. Improved control margins will be achieved by avoiding winds in the critical relative wind azimuth areas (figure 4-1).

4-5-A. HOVER CEILING - IGE

OGE at all pressure altitude and outside air temperature conditions with heater on or off at maximum continuous engine power. CAUTION

OGE HOVER OPERATION MAY RESULT IN VIOLATION OF HEIGHT-VELOCITY LIMITATIONS.

AREA A (unshaded area) as shown on the hover ceiling charts presents hover performance for which satisfactory cyclic have been control and directional demonstrated in relative winds of 35 knots from any direction at or below 3000 feet HD. Improved control margins will be achieved by avoiding winds in the critical relative wind azimuth areas (figure 4-1).

AREA B (shaded area) as shown on hover ceiling charts presents additional hover performance which can be achieved in calm winds or winds outside the critical relative wind azimuth areas. NOTE

The Hover Celling chart (figure 4-2) provides the maximum allowable gross weights for hovering IGE at all pressure altitude and outside air temperature conditions with heater on or off at maximum continuous engine power. Conversely, the hover ceiling altitude can be determined for any given gross weight.

Tail rotor or cyclic control margin may preclude operation in AREA B of the hover ceiling charts when the relative wind is in the respective critical wind azimuth area.

4-5-B. HOVER CEILING - OGE

The Hover Ceiling charts (figure 4-3) provide maximum weights for hovering 3


FAA APPROVED

NOTES: 1. Pedal critical wind azimuth-hovering with the relative wind within these azimuth angles can result in the following: a. Inability to maintain heading due to large left pedal requirements for certain wind velocities. b. Reduction of available left pedal control with a directional AFCS hardover. 2. Longitudinal cyclic criticalwind azimuth aft cyclic may be limited with longitudinal AFCS hardover.

270

SEENOTE 2


SEENOTE

180

IGE CRITICAL RELATIVE WIND AZIMUTH

Figure 4-1. Critical relative wind azimuths

412FMS-41.3-4-1


FAA APPROVED


SKID HEIGHT 4 FEET HEATER ON OR OFF -400 TO 520C



-CY

…

/

-/

EIGHTLI'IT

1A\\

I

if

-40

-20

0 C2 20 ~

OAT

-

~

C

40 '\-

60H

8

9

10

11

12

GROSS WEIGHT - LBS x 1000

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

0-~~~~

412FS-41.3.4-2

~~Figure 4-2. Hover ceiling IGE

^~~

Rev.2 5 \\^^^^^^

SEP 5 2000

~SEP 05 Z008

.~v.2

5


FAA APPROVED

HOVER CEILING


MAXIMUM CONTINUOUS POWER SKID HEIGHT 60 FT. ENGINE RPM 100%/ HEATER OFF 0 TO 52°C GENERATOR 150 AMPS(EA.) CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

14,000 FT HD UMIT

,o% 1

1\

:<;^~ y

0

10

20 30 OAT - C

ÂREA Ai

40

50

8 9 10 11 12 GROSS WEIGHT - LBS x 1000 I 'I Im i Ii I 3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

I

412FS-41.3-4-3-1

Figure 4-3.

6

Rev. 2

SEP 05 2000

Hover ceiling OGE (Sheet 1 of 4)

41


BHT-412-FMS-41.3

HOVER CEILING

*^~~B~~OUT

OF GROUND EFFECT


SKID HEIGHT 60 FT. HEATER OFF -40 TO 0° C


14,000 FT HD LIMIT

.

-40

-30

L

-20 -10 OAT - C .

-

0

,

°-2)

t I-~;I_\/ -

8 9 10 11 12 GROSS WEIGHT - LBS x 1000 I I I I I 3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT -

KG X 1000

I_____________________________________ _____________412FS-41.3-4-3-2 Figure 4-3.


SEP 052000

Rev. 2

7


BHT-412-FMS-41.3

HOVER CEILING





14,000 FT HD UMIT

0

10 OAT -

12 11 10 9 8 20' GROSS WEIGHT - LBS x 1000 °C I, I I I I I [ 5.4 4.6 5.0 3.4 3.8 4.2 GROSS WEIGHT - KG X 1000

412FS41.3-4-3

Figure 4-3.

8

Rev. 2

IP

5 20O


T N I R P N O R I O T F C U T O OD N R E P E E R R _,=C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-41.3

HOVER CEILING


MAXIMUM CONTINUOUS POWER SKID HEIGHT 60 FT. ENGINE RPM 100% HEATER ON GENERATOR 150 AMPS(EA.) -40 TO 0°C CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

14.000 FT HD LIMIT

AEAI

-

-40

-30

-20

OAT -

/', __/

AI!

E

-10

0

°C

B-AFEA.A

8

9

OAT

10

~-

11

12

GROSS WEIGHT - LBS x 1000 I I II I I I I I [` ]

3.4

3.8

4.2

4.6

GROSS WEIGHT -

5.0

5.4

KG X 1000

412FS-41.3-43-4

Figure 4-3. Hover ceiling OGE (Sheet 4 of 4)

SEP 05 20Q

Rev. 2

9/10



SPECIAL MISSION SEARCHAND RESCUE CONFIGURATION 412-899-052

BHT-36020 -

36086

BHT-36087 AND SUB CERTIFIED 8 MARCH1993

This supplement shall be attached to Model 412&412EP Flight Manual when SPECIAL MISSION SEARCH AND RESCUE CONFIGURATIONkit has been installed.

Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

INC BELL HELICOPTER ANDBELLHELICOPTER TEXTRON INC A DIVISION OF TEXTRON CANADA LT. ALLRIGHTS RESERVED

I

of TextronInc A Subsidiary

NOTICE COPYRIGHT

REISSUE

-

2

FEBRUARY

1995


FAA APPROVED

NOTICEPAGE





BHT-412-FMS-43.3 AND 43.4

LOG OF REVISIONS

Original ........... Revision..........1

0 ............... ..............

8 MAR 93 22 APR 93

Reissue ........... Reissue ...........

0 .............. 0 ...............

16 SEP 93 2 FEB 95

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

T itle .................................... NP ...................................... A - B .................................. i/ii ....................................... 1 - 4 ...................................

REVISION NO.

PAGE

5/6 ......................................

0 0 0 0 0

0

MANUFACTURER'S DATA

7-14 .................................. 15/16 ...................................

0 0

NOTE



FAA APPROVED


Original ........... Revision ..........

0...............8 1 ..............

MAR 93 22 APR 93

APPROVED:

MANAGER


Reissue ........... Reissue...........0

0............16 ...............

SEP 93 2 FEB 95


FAA APPROVED

GENERAL INFORMATION

The Special Mission Search And Rescue Configuration helicopter is equipped with the following optional equipment kits: Rotor brake

3-Place litter

Dual controls

Radar altimeter

Copilot clock

Cargo hook

FAA/IFR

ADF (KDF806)

VHF comm 2 (KTR-908)

NAV #1 VOR/LOCwith glideslope (KNR-634)

NAV #2 VOR/LOCwith glideslope (KNR-634)Transponder (KXP-756) Emergency floats (reservoir/nose)

Emergencyfloats provisions

Emergency floats (L/G with floats)

Auxiliary fuel provisions (168 gallons)

Auxiliary fuel, left hand (84 gallons)

Increased generator capacity

Increased horsepower transmission

Copilot instruments

Standby attitude indicator

HF communication (provisions)

NAT 15 Station ICS

DF931-1and VHF/UHFAM guard receivers

Radio operator station (provisions)

HF cypher (provisions)

MS-AVCIDIAdigital FM radio (provisions)

Tow rings

High tiedown fittings

DME670

#2 Radio altimeter

EFISdisplay (EDZ-705)

Primus 700 weather radar

RACALsystem 91 Doppler

Data nav III

Rappeling rings

SX-16C nightsun

Loudhailer, AEM 400

Infrared scanner (FLIR Safire)

ELT 100 HM (provisions)

External hoist (Lucas Western, 42325-1), Three 10A 28VDCoutlets

Dual digital automatic flight control system, Search and Rescue (SAR) Ground handling cable and claw assembly GNS-X NO LORANwith VLF/OMEGAand GPS and video output When above equipment is installed on helicopters other than Special Mission Search And Rescue Configuration, refer to

appropriate Flight Manual Supplement (FMS) for applicable limitations and operations.

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-43.3 AND BHT-412-FMS-43.4

1-1. INTRODUCTION

1-5-A-2.

Limitations specified in BHT-412-FM-3, BHT-412-FM-4 and supplements remain applicable. Operation of SPZ-7600system flight crew of does not relieve responsibility to monitor and comply with all helicopter limitations.


For Search and Rescue, Dual Digital Automatic Flight Control System Limitations, refer to BHT-412-FMS-39.3or BHT-412-FMS-39.4.

Actual weight change shall be determined has been properly after helicopter configured and ballast readjusted, if necessary, to maintain CG limits within limits shown on Gross weight center of gravity charts in BHT-412-FM-3and BHT412-FM-4.

For External Hoist Limitations, refer to Premier Aviation Flight Manual in Supplement for hoist installed accordance with STC SH7860SW.


Special Mission Search And Rescue Configuration helicopters are certified for IFR/VFR operations under day or night non-icing conditions.

1-5. CONFIGURATION

1-5-A. REQUIREDEQUIPMENT 1-5-A-1. NIGHTSUN SEARCHLIGHT

Do not use nightsun searchlight below 50 feet AGL. Do not use nightsun searchlight in foggy or cloudy conditions. Operation of nightsun searchlight is restricted to copilot or operator position.

1-6-A. WEIGHT

1-6-B. CENTEROF GRAVITY

For flights with no internal loading and any optional equipment kits removed, a weight and balance check must be calculated to determine a new minimum cockpit weight required to maintain helicopter within aft CG limits.

For flights with more than seven (7) passengers or crewmembers in forward and aft facing seats and all optional equipment kits installed, a weight and balance must be calculated to determine if any baggage compartment weight is required to maintain helicopter within forward CG limits.

1-6-C. LATERALAND LONGITUDINALCENTEROF GRAVITY

Maximum hoist load shall be adjusted to remain within helicopter lateral and longitudinal CG limits. Refer to hoist 1

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I Section 2 R W O E I F V NLY O BHT-412-FMS-43.3 AND BHT-412-FMS-43.4 loading schedule in Premier Aviation Flight Manual Supplement for hoist installed in accordance with STC SH7860SW.

FAA APPROVED

1-20. INSTRUMENT MARKINGSAND PLACARDS Refer to figure 1-1.

CAUTION

DO NOT USE NIGHTSUN SEARCHLIGHT BELOW 50 FT AGL OR IN FOG CONDITIONS. FOR TAXI TAKEOFF AND LANDING SEARCHLIGHT MUST BE IN HORIZONTAL OR UP STOWED POSITION. MONITOR LOADMETER WHEN USING NIGHTSUN SEARCHLIGHT. LOCATED ON INSTRUMENT PANEL

WITH FLIR TURNED ON. VHF COMM NO. 2 IS UNRELIABLE WHEN TUNED TO 128.00 MHz.

UHF/MARINE GUARD RECEIVER FUNCTION IS UNRELIABLE DURING RADIO TRANSMISSION OR ACTUATION OF NIGHTSUN GIMBAL. LOCATED ON PEDESTAL

BHT-412-FS-42.3-1-1

Figure 1-1.

2-1. INTRODUCTION

Normal Procedures specified in BHT-412FM-3, BHT-412-FM-4 and supplements remain applicable.

For Search and Rescue, Dual Digital Automatic Flight Control System Normal Procedures, refer to BHT-412-FMS-39.3or BHT-412-FMS-39.4.

Decal

For External Hoist Normal Procedures, refer to Premier Aviation Flight Manual in for hoist installed Supplement accordance with STC SH7860SW.

For Forward Looking Infrared Radar (FLIR) Normal Procedures, refer to TM 85000750 (FLIR Systems Inc.) Operator's Manual.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3 AND BHT-412-FMS-43.4

FAA APPROVED

2-3. PREFLIGHTCHECK

2-3-A. NIGHTSUN SEARCHLIGHT

Nightsun searchlight - Security and wiring. Lens for cleanliness. Stowed

position.

SEARCHLIGHT IS OPERATING. NIGHTSUN SEARCHLIGHT SHALL BE TURNED OFF WHEN ENTERING CLOUDS OR FOG.

NIGHTSUN FREE/STOW switch - Press and release, check FREE segment illuminated. START switch -


CAUTION

NIGHTSUNcircuit breaker - In. Loudhailer POWER switch -

HOLDING SWITCH IN START AFTER IGNITION MAY DAMAGE EQUIPMENT.

Off.

2-6. SYSTEMSCHECK

NIGHTSUN FREE/STOW switch STOW segment illuminated.

Check

START switch - START. Hold in START approximately 5 seconds or until ignition has occurred. Allow switch to return to ON.

NIGHTSUN FREE/STOW switch - Press and release, check FREE segment illuminated. NIGHTSUN FREE/STOW switch - Press and release, check STOW segment illuminated.

GROUND

CARE DURING

OPERATION

WARNING

DO NOT AIM BEAM TOWARD OTHER AIRCRAFT OR VEHICLES BECAUSE OF TEMPORARY BLINDING EFFECT.

UP/DOWN/LEFT/RT aim light.

WARNING

USE EXTREME

ON.

OF

LOUDHAILER TO PREVENT INJURY TO PERSONNEL. GROUND SUPPORT PERSONNEL IN VICINITY OF HELICOPTER SHOULD WEAR PROTECTIVE

HEARINGDEVICES.

FOCUS switch -

ICS control panel HAIL. WAILSIREN/YELP

Position

to

Press to focus light.

Loudhailer POWER switch -

VOL switch -


switch -

ON.

Rotate switch to AUX/

switch -

As desired.

As desired.

Loudhailer POWER switch - Off, when loudhailer operation is completed.

WARNING

MAGNETIC COMPASS MAY BE UNRELIABLE WHEN NIGHTSUN 3

T N I R P N O Section3R I O T F C U T O OD N R E P E E R R Section 4 C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3AND BHT-412-FMS-43.4


FAA APPROVED

extinguished for approximately 30 seconds until searchlight returns to stowed position. STOWsegment then illuminates.

NIGHTSUN FREE/STOW switch - Press and release. FREE/STOW lights remain

Emergency and Malfunction Procedures specified in BHT-412-FM-3,BHT-412-FM-4 and supplements remain applicable. For Search and Rescue, Dual Digital Automatic Flight Control System Emergency and Malfunction Procedures, refer to BHT-412-FMS-39.3 or BHT-412FMS-39.4.

4-7. CLIMB AND DESCENT

Rate of climb performance derived from Twin Engine Rate of Climb Charts, Maximum Continuous Power and Single Engine Rate of Climb Charts, Maximum Continuous Power shown in BHT-412-FM-3 or BHT-412-FM-4shall be reduced by 110 feet per minute. Additional rate of climb loss may result from Increased Generator Capacity Operation (refer to BHT-412-FMS-40).

For External Hoist Emergency and Malfunction Procedures, refer to Premier Aviation Flight Manual Supplement for hoist installed in accordance with STC SH7860SW.

For Forward Looking Infrared Radar (FLIR) Malfunction Procedures, refer to TM 85000750 (FLIR Systems Inc.) Operator's Manual.

4-8. AIRSPEED CALIBRATION

As a result of increased drag caused by installation of optional kits, reduce airspeeds derived from Fuel Flow vs Airspeed Charts in Section 3 of BHT-412MD-3or BHT-412-MD-4by 10 knots.

Performance data shown In BHT-412-FM-3, BHT-412-FM-4and supplements (except as stated above) remain applicable.

For Search and Rescue, Dual Digital Automatic Flight Control System Performance Data, refer to BHT-412-FMS39.3 or BHT-412-FMS-39.4.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3AND BHT-412-FMS-43.4

FAA APPROVED

For External Hoist Performance data, refer to Premier Aviation Flight Manual Supplement for hoist installed in accordance with STC SH7860SW.

For Increased Generator Capacity Performance data, refer to BHT-412-FMS 40.

4-10. NOISE LEVELS

4-10-A. CERTIFICATED FAR PART 36 STAGE 2 NOISE LEVELS

This aircraft is certified as a Stage 2 helicopter as prescribed in FAR Part 36, Subpart H, for gross weights up to, and including, the certified maximum takeoff and landing weight of 11,900 pounds (5398 kilograms). There are no operating limitations in meeting the takeoff, flyover, or approach noise requirements. The following noise levels comply with FAR Part 36, Appendix H, Stage 2 noise level requirements. These levels were obtained by analysis of approved data from noise tests conducted under provisions of FAR Part 36, Amendment 3614. The certified noise levels are: FLIGHT CONDITION

EPNL (EPN dB)

Takeoff Flyover Approach

93.7 92.9 95.3

NOTE

No determination has been made by the Federal Aviation Administration that the noise levels of this aircraft are, or should be, acceptable or unacceptable for operations at, into, or out of any airport.

VH is defined as the airspeed in level flight obtained using minimum specification engine torque corresponding to maximum continuous power available for sea level, 25°C (77°F) ambient conditions at the relevant maximum certified weight. The value of VHthus defined for this aircraft is 112 KTAS.

4-10-B. SUPPLEMENTAL ICAO ANNEX 16, CHAPTER 8 NOISE LEVEL INFORMATION

The test and analysis procedures used to obtain these noise levels are essentially equivalent to those required by the International Civil Aviation Organization (ICAO) in Annex 16, Volume 1, Chapter 8. ICAO Annex 16, Volume 1, Chapter 8 only after is applicable approval endorsement by the Civil Aviation Authority of the country of aircraft registration.

5/6

T Section 5 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

5-5.

COCKPIT AND CABIN

LOADING

5-5-A. STATIONDIAGRAMS


5-8. SAMPLELOADING PROBLEM

A sample loading problem is presented in both English and Metric.

Figure 1-1 presents fuselage station (FS) and buttock line (BL) data to aid in weight and balance computation.

5-5-B. PERSONNELLOADING

Table 1-1 presents personnel weights and moments in both English and Metric units to aid in weight and balance computation.

7

T N I R P N O R I O T I F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3 AND BHT-412-FMS-43.4

MANUFACTURER'S DATA

NOTE Station 0 (reference datum) is located 20 inches (508 millimeters) aft of the most forward point of the cabin nose.

FUSELAGE STATIONS

243

23

FORWARD FACING

REF

DATUM

COPILOTOR PASSENGERSEAT

Figure 5-1. Station diagrams (Sheet 1 of 2) (Forward row seats - forward facing)



NOTE Station 0 (reference datum) is located 20 inches (508 millimeters) aft of the most forward point of the cabin nose.

FUSELAGE STATIONS

23

AFT FACING

REF DATUM

Figure 5-1. Station diagrams (Sheet 2 of 2) (Forward row seats - aft facing)

9

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3 AND BHT-412-FMS-43.4 Table 5-1.

MANUFACTURER'S DATA

Personnel loading tables

PILOT AND PASSENGER TABLE OF MOMENTS (IN-LB)

Weight (LB) 100 110 120 130 140 150 160 170 180 190 200 210 220

Pilot and

Passenger (1-Man Seats

Copilot FS 47 4700 5170 5640 6110 6580 7050 7520 7990 8460 8930 9400 9870 10340

Facing Aft) FS 79.7 7970 8767 9564 10361 11158 11955 12752 13549 14346 15143 15940 16737 17534

Passenger (1-Man Seats Facing Fwd) FS 85.7 8570 9427 10284 11141 11998 12855 13712 14569 15426 16283 17140 17997 18854

Passenger

PASSENGER

(5-Man Seat Facing Fwd) FS 114.7 11470 12617 13764 14911 16058 17205 18352 19499 20646 21793 22940 24087 25234

FACING OUTBOARD

Fwd Seat FS 139.7 13970 15367 16764 18161 19558 20955 22352 23749 25146 26543 27940 29337 30734

Aft Seat FS 156.8 15680 17248 18816 20384 21952 23520 25088 26656 28224 29792 31360 32928 34496

*Left Forward Seat

PILOT AND PASSENGER TABLE OF MOMENTS (KG * MM)

Weight (KG) 50 55 60 65 70 75 77.1 80 85 90 95 100 105 110

Pilot and Copilot* 1194 MM 597.0 656.7 716.4 776.1 835.8 895.5 920.6 955.2 1014.9 1074.6 1134.3 1194.0 1253.7 1313.4

Passenger (1-Man Seats Facing Aft) 2024 MM 1012.0 1113.2 1214.4 1315.6 1416.8 1518.0 1560.5 1619.2 1720.4 1821.6 1922.8 2024.0 2125.2 2226.4

100 Passenger Passenger (1-Man Seats (5-Man Seat Facing Fwd) Facing Fwd) 2177 MM 2913 MM 1088.5 1456.5 1197.4 1602.2 1306.2 1747.8 1415.1 1893.5 1523.9 2039.1 1632.8 2184.8 1678.5 2245.9 1741.6 2330.4 1850.5 2476.1 1959.3 2621.7 2068.2 2767.4 2177.0 2913.0 2285.9 3058.8 2394.7 3204.3

PASSENGER

FACING OUTBOARD Fwd Seat Aft Seat 3548 MM 3983 MM 1774.0 1991.5 1951.4 2190.7 2128.8 2389.8 2306.2 2589.0 2483.6 2788.1 2661.0 2987.3 2735.5 3070.9 2838.4 3186.4 3015.8 3385.6 3193.2 3584.7 3370.6 3783.9 3548.0 3983.0 3725.4 4182.2 3902.8 4381.3

*Left Forward Seat (TABLE

I.D.

911317)


SAMPLE


LOADING PROBLEM

(ENGLISH)

The helicopter is chartered to transport nine passengers and 180 pounds of baggage on a trip that will require approximately 235 U.S. gallons of fuel one way. The helicopter will be refueled to full fuel and both pilots will make the return trip. Determine extreme CG conditions for both flights.

OUTBOUNDFLIGHT

LONGITUDINAL

WEIGHT

Weight Empty

8223

+Oil +Crew (2) +Passenger, Aft Row (5) +Passenger, Aft Facing (4) +Baggage +Fuel to maximum GW

25 340 850 680 180 1602

CG

140.85

114.7 79.7 261.0 150.4

LATERAL

MOMENT 1158196

4146 15980 97495 54196 46980 240941

CG

MOMENT

+0.6

+5007

0 0 0

0 0 0 +17 0 -9260

0 -5.78

GW and most aft CG -Fuel to critical forward CG

11900 -1205

136.0

1617934 -185401

-0.4

-4236 +9260

Most forward CG

10695

133.9

1432533

+0.5

+5024

RETURN FLIGHT

LONGITUDINAL

WEIGHT

Weight Empty

8223

+Oil +Crew (2) +Fuel, including auxiliary fuel

25 340 2802

GW and most aft CG -Fuel to critical forward CG Most forward

CG

CG

140.85

LATERAL

MOMENT

1158196

4146 15980 422822

CG

+0.6

MOMENT +5007

0 0 -6.03

0 0 -16896

11390 -2405

140.6

1601144 -367282

-1.0

-11889 +16896

8985

137.3

1233862

+0.6

+5007

11

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3 AND BHT-412-FMS-43.4

MANUFACTURER'S DATA

SAMPLE LOADING PROBLEM

(METRIC)

The helicopter is chartered to transport nine passengers and 80 kilograms of baggage for a trip that will require approximately 890 liters of fuel one way. The helicopter will be refueled to full fuel and both pilots will make the return trip. Determine extreme CG conditions for both flights. OUTBOUND FLIGHT

LONGITUDINAL

Weight Empty

LATERAL

WEIGHT (KG)

CG (MM)

MOMENT KG*MM 100

CG (MM)

MOMENT KG*MM 100

3729.9

3578

133455.8

+15

+559.5

+Oil +Crew (2) +Passenger, Aft row (5) +Passengers, Aft facing (4) +Baggage +Fuel to maximum GW

11.3 154.2 385.6 308.4 80.0 728.3

1194 2913 2024 6629 3822

485.3 1841.1 11232.5 6242.0 5303.2 27835.6

0 0 0 +1 0 -147

0 0 0 +2.0 0 -1070.6


5397.7 -548.4

3453

186395.5 -21443.8

-9

-509.1 +1070.6

Most forward CG

4849.3

3402

164951.7

+12

+561.5

RETURN FLIGHT

LONGITUDINAL

LATERAL

WEIGHT (KG)

CG (MM)

MOMENT KG*MM 100

CG (MM)

Weight

3729.9

3578

133455.8

+15

+559.5

+Oil +Crew (2) +Fuel including auxiliary fuel

11.3 154.2 1271.0

1194 3834

485.3 1841.1 48730.1

0 0 -153

0 0 -1944.6


5166.4 -1091.1

3571

184512.3 -42338.3

-27

-1385.1 +1944.6

Most forward

4075.3

3489

142174.0

+14

+559.5

CG

MOMENT KG*MM 100

(TABLE

I.D.

910527)

T Section1 N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

AND BHT-412-FMS-43.4 BHT-412-FMS-43.3

1-1. NIGHTSUN SEARCHLIGHT

Nightsun searchlight operates on 28 VDC power obtained from non-essential bus through NIGHTSUNcircuit breaker ( amp) located in overhead console. A 70 amp remote control circuit breaker is also installed below roof, aft of overhead console.

( LOCATED ON PEDESTAL )

1-1-A. CONTROLS 1-1-A-1.

FREE/STOW SWITCH

NIGHTSUNFREE/STOWswitch (figure 1-1) is a pushbutton press-to-display condition switch located on miscellaneous control panel. When helicopter power is on and NIGHTSUN circuit breaker in, STOW (green) segment is illuminated. When switch is pressed and released, FREE (amber) segment illuminates and nightsun searchlight may be started and operated. After nightsun searchlight is turned off, for are extinguished segments approximately 30 seconds until nightsun returns to searchlight automatically Once nightsun stowed position. searchlight is stowed, STOW segment is illuminated. 1-1-A-2.

Figure 1-1. Miscellaneous control panel

START switch (figure 1-2) is a threeposition toggle switch located on control box. Switch is springloaded to return to ON from START after switch is released. When NIGHTSUN FREE/STOW switch (figure 1-1) FREE segment is illuminated, placing START switch to ON, then holding in START (for approximately 5 seconds) illuminates the nightsun searchlight. When operation of nightsun searchlight is completed, positioning START switch to OFF removes power from nightsun searchlight.

START SWITCH CAUTION

HOLDING SWITCH IN START POSITION AFTER IGNITION MAY DAMAGEEQUIPMENT. 13

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-43.3AND BHT-412-FMS-43.4

MANUFACTURER'SDATA

released, switch will return to a neutral position. 1-1-A-4.

FOCUS SWITCH

FOCUS switch (figure 1-2) is located on control box and is pressed to concentrate light beam when nightsun searchlight is in operation.

DOWN

LEFT

RT

1-1-B.

DF 931 CONTROL

UP

1-1-B-1.

Figure 1-2. Control box

1-1-A-3.

UP/DOWN/LEFT/RT SWITCH WARNING

DO NOT AIM BEAM TOWARD OTHER AIRCRAFT OR VEHICLES BECAUSE OF TEMPORARY BLINDING EFFECT.

UP/DOWNILEFT/RTswitch (figure 1-2) is a five-position toggle switch located on control box. When nightsun searchlight Is illuminated, it can be aimed by using this switch. When light is aimed and switch is

DF FUNCTION SWITCH

DF function switch (figure 1-1) OFF/MAR/ UHF/VHF is used to select MAR guard, UHF guard, or VHF (COMM 2) receivers. By placing DF function switch in any position except OFF, both guard receivers will monitor their selected frequency. 1-1-B-2.

AUX/MAIN SWITCH

When AUX/MAIN switch is in MAIN, MAR receiver monitors marine guard channel (156.800 MHz) and UHF receiver monitors UHF guard channel (243.000MHz). When AUX/MAINswitch Is in AUX, MAR receiver is tuned to 161.200MHz and UHF receiver is tuned to 245.100 MHz. AUX position is used only for maintenance.

VHF DF position Is controlled by frequency selected on VHF comm 2 control head. To display selected DF receiver on EFIS indicator, place EFIS mode switch to DF. If DF audio is desired from guard receivers, place ADF/DF switch on ICS panel to DF.



1-2. ELECTRICALSYSTEM

1-2-A. DC ELECTRICAL SYSTEM

Table 1-1. DC electrical system bus loading (Nonessential buses)

A

A



NAV-COMMSYSTEMS CABIN ICS RADIO OPERATORPANEL RADIO OPERATORLT

NAV-COMMSYSTEMS LR NAV MGMT LR NAV RCVR FM DG FM HF CIPHER DF

MISCELLANEOUSSYSTEMS NIGHTSUN AUX OUTLETSRH AUX OUTLETSCTR AUX OUTLETSLH

MISCELLANEOUSSYSTEMS FLIR SYS FLIR MONITOR HAIL

NOTES:

A

Automatically sheds with one generator inoperative, but can be reenergized by placing NON-ESTL BUS switch in MANUAL position. Automatically battery only.

Table 1-2.

sheds for 30 and 90 minute flight with electrical power supplied by

DC electrical

system bus loading (Battery bus)

DC BATTERY BUS

MISCELLANEOUS SYSTEMS HOIST POWER

HOIST CONTROL

15/16

BHT-412-FMS-44.3AND -44.4

T N I R P N ROTORCRAFT O R I FLIGHT MANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODELS

SUPPLEMENT

AIRLINEPASSENGERSEATING 412-706-028

BHT 36020 - 36086 AND BHT 36087 AND SUB CERTIFIED 18 JUNE 1993

This supplement shall be attached to Model 412 Flight Manual (BHT-412-FM-3)or Model 412EP Flight Manual (BHT-412-FM-4) Flight Manual when the 412-706-028 Airline Passenger Seating Kit has been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, or other applicable supplements, consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1999 BELL HELICOPTER TEXTRON INC. ANDBELLHELICOPTER TEXTRON INC. A DIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

A Subsidiary ofTextron Inc. POST OFFICE BOX 482


05 OCTOBER 1994 REVISION1 - 13 JULY 1999

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-44.3AND -44.4

FAA APPROVED

NOTICEPAGE


Manufacturer's Data portion of this supplement is proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation Is forbidden without prior written authorization from Bell Helicopter Textron, Inc.


NP

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-44.3 AND -44.4

LOGOF REVISIONS


0..............18

JUN 93

........................ Reissue . Revision .......... 1 ..............

05 OCT 94 13 JUL 99

LOGOF PAGES

REVISION NO.

PAGE

FLIGHTMANUAL

T itle .................................... N P ...................................... A - B .................................. 1 .......................................

REVISION NO.

PAGE

0

2 .......................................

1 0 1 1

MANUFACTURER'SDATA

0 0

3- 6 ................................... 7/8 ......................................

NOTE

Revised text Is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. Rev.1

A

BHT-412-FMS-44.3AND -44.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOGOF FAAAPPROVED REVISIONS

Original ........... 0 .............. Reissue...........0 .............. Revision.......... ..............

18 JUN 93 05 OCT 94 13 JUL 99

APPROVED:

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AVIATIONADMINISTRATION FT.WORTH,TX 76193-0170

B

Rev.1

DATE:

BHT-412-FMS-44.3AND -44.4

FAA APPROVED

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S Section2N G O N I R W O E I F V NLY O 1-5. CONFIGURATION

1-5-C. DOORSOPEN/REMOVED

Seat backs and cushions for two outboard facing seats on each side of transmission housing, If unoccupied, shall be removed prior to operations with sliding doors open or removed.

Actual weight change shall be determined after kit is installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits.

1-7. AIRSPEED V

with doors

KIAS.

open or removed is 100



No change from basic manual. Rev.

1

FAA APPROVED

BHT-412-FMS-44.3AND -44.4

T N Section4 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O No change from basic manual.

T N I Section 5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

5-3. GROSSWEIGHT CENTER OF GRAVITY

Figure 1-1 presents fuselage station and buttock line data to aid in weight and balance computations.

BHT-412-FMS-44.3 AND -44.4

5-8. SAMPLE LOADING PROBLEM

A sample loading problem is provided in both English and Metric units.

5-5. COCKPITAND CABIN LOADING

Refer to table 1-1 for personnel weights and moments in both English and Metric units.

3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-44.3AND -44.4

NOTE

Station 0 (reference datum) is located 20 inches (508 millimeters) aft of most forward point of cabin nose.

FUSELAGE STATIONS

23

166

REFERENCE DATUM

STANDARDFRONT ROW CONFIGURATION

FUSELAGE

REFERENCE DATUM

Figure1-1. Stationdiagram

243


BHT-412-FMS-44.3 AND -44.4

Table 5-1.

Crew and passenger table of moments English (IN · LB)

Weight (LB) 100 110 120 130 140 150 160 170 180 190 200 210 220

Pilot and Copilot* FS 47 4700 5170 5640 6110 6580 7050 7520 7990 8460 8930 9400 9870 10340

Passenger (Forward 4-Man Seat) FS 85.7 8570 9427 10284 11141 11998 12855 13712 14569 15426 16283 17140 17997 18854

Passenger (Aft 5-Man Seat) FS 114.7 11470 12617 13764 14911 16058 17205 18352 19499 20646 21793 22940 24087 25234

Passenger Facing Outboard Fwd Seat Aft Seat FS 139.7 FS 156.8 13970 15680 15367 17248 16764 18816 18161 20384 19558 21952 20955 23520 22352 25088 23749 26656 25146 28224 26543 29792 27940 31360 29337 32928 30734 34496

*Left Forward Seat

Metric (KG * MM/100)

Weight (KG) 50 55 60 65 70 75 77 .1 80 85 90 95 100 105 110

Pilot and Copilot* 1194 MM 597.0 656.7 716.4 776.1 835.8 895.5 920.6 955.2 1014.9 1074.6 1134.3 1194.0 1253.7 1313.4

Passenger (Forward 4-Man Seat) 2177 MM 1088.5 1197.4 1306.2 1415.1 1523.9 1632.8 1678.5 1741.6 1850.5 1959.3 2068.2 2177.0 2285.9 2394.7

Passenger (Aft 5-Man Seat) 2913 MM 1456.5 1602.2 1747.8 1893.5 2039.1 2184.8 2245.9 2330.4 2476.1 2621.7 2767.4 2913.0 3058.7 3204.3

Passenger Facing Outboard Fwd Seat Aft Seat 3548 MM 3983 MM 1774.0 1991.5 1951.4 2190.7 2128.8 2389.8 2306.2 2589.0 2483.6 2788.1 2661.0 2987.3 2735.5 3070.9 2838.4 3186.4 3015.8 3385.6 3193.2 3584.7 3370.6 3783.9 3548.0 3983.0 3725.4 4182.2 3902.8 4381.3

*Left Forward Seat

(TABLE

I.D. 910522)

5


BHT-412-FMS-44.3 AND -44.4

SAMPLE

LOADING PROBLEM (ENGLISH)

The helicopter is chartered to transport nine passengers and 180 pounds of baggage on a trip that will require approximately 330 U.S. gallons of fuel one way. The helicopter will be refueled to full fuel and both pilots will make the return trip. Determine extreme CG conditions for both flights.

OUTBOUNDFLIGHT

LATERAL

LONGITUDINAL

MOMENT (IN-LB)

143.2

1031040

47.0 114.7 85.7 261.0 151.5

4146 15980 97495 58276 46980 340421

0 0 0 -5.25 0 -0.4

0 0 -3570 0 -854

11522 -1850

138.4

1594338 -284881

-0.3 -0.5

-2984 +854

9672

135.4

1309457

-0.2

-2130

Weight Empty

7200

+Oil +Crew (2) +Passenger, Aft Row (5) +Passenger, Forward Row (4) +Baggage +Fuel (full)

25 340 850 680 180 2247

GW and most aft CG -Fuel to critical forward CG Most forward CG

CG (IN)

MOMENT (IN-LB)

CG (IN)

WEIGHT (LB)

+02

+1440 0

RETURN FLIGHT

LATERAL

LONGITUDINAL

WEIGHT (LB)

CG (IN)

MOMENT (IN-LB)

CG (IN)

MOMENT (IN-LB)

143.2

1031040

+0.2

47.0 151.5

4146 15980 340421

0 0

-0.4

-854

+1440

Weight Empty

7200

+011 +Crew (2) +Fuel (full)

25 340 2247


9812 -1850

141.8

1391587 -284881

+0.1 -0.5

+586 +854

Most forward CG

7962

139.0

1106706

+0.2

+1440

0 0

From Gross weight center of gravity chart in Section 1 of basic flight manual, it can be determined that CG will remain within limits for outbound and return flights. (TABLE

6

I.D. 910520)


BHT-412-FMS-44.3 AND -44.4

SAMPLE LOADING PROBLEM (METRIC)

The helicopter is chartered to transport nine passengers and 80 kilograms of baggage for a trip that will require approximately 1250 liters of fuel one way. The helicopter will be refueled to full fuel and both pilots will make the return trip. Determine extreme CG conditions for both flights. OUTBOUND FLIGHT

LONGITUDINAL

WEIGHT (KG)

LATERAL

CG (MM)

MOMENT ( KG*MM 100

CG (MM)

3637

118777.1

-+5

+163.3

1194 2913 2177 6629 3848

485.3 1841.1 11232.5 6713.9 5303.2 39218.8

0 0 0 -133 0 -10

0 0 0 -410.2

(

MOMENT KG )MM 100

Weight Empty

3265.8

+Oil +Crew (2) +Passenger, Aft Row (5) +Passengers, Forward Row (4) +Baggage +Fuel (full)

11.3 154.2 385.6 308.4 80.0 1019.2


5224.5 -839.3

3514

183571.9 -32827.0

-7 -12

-348.8 +101.9

Most forward CG

4385.2

3438

150744.9

-6

-246.9

0

-101.9

RETURN FLIGHT

LONGITUDINAL

LATERAL

WEIGHT (KG)

CG (MM)

MOMENT KG-MM 100

MOMENT ( KG-MM)

(MM)

Weight

3265.8

3637

118777.1

+5

+163.3

+Oil +Crew (2) +Fuel (full)

11.3 154.2 1019.2

1194 3848

485.3 1841.1 39218.8

0 0 -10

0 0 -101.9


4450.5 -839.3

3602

160322.3 -32827.0

+1 -12

+61.4 +101.9

Most forward CG

3611.2

3531

127495.3

+5

+163.3

CG

100

From Gross weight center of gravity chart in Section 1 of basic flight manual, it can be determined that CG will remain within limits for outbound and return flights. (TABLE

I.D. 910521)

7/8

BHT-412-FMS-45.3 & 45.4

T N I R P N ROTORCRAFT O R I MANUAL FLIGHT O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Bell 412/412 EP MODELS

SUPPLEMENT

INCREASED MAXIMUM CONTINUOUS POWER KIT WITH PT6T-3D ENGINE (412-706-029)

S/N 36020 - 36080 AND S/N 36081 AND SUB CERTIFIED 27 AUGUST 1993

This supplement shall be attached to Flight Manual BHT-412-FM-3 and BHT-412-FM-4 when Increased Maximum Continuous Power Kit (412-706-029) has been installed.

* Helicopters S/N 36081 through 36086 are equipped with PT6T-3D engines. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, or other applicable supplements, consult basic Flight Manual.

Bell HelicopterTEXTRON NOTICE COPYRIGHT COPYRIGHT 2000 INC. TEXTRON BELL HELICOPTER ANDBELLHELICOPTER TEXTRON INC. A DIVISION OFTEXTRON CANADA LTD ALLRIGHTSRESERVED

A Subsidiaryof TextronInc. POST OFFICE BOX 482 FORT WORTH, TEXAS 76101

23 JUNE REVISION 3-5

1994

SEPTEMBER 2000

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-45.3 &45.4

FAA APPROVED

NOTICE PAGE


NP

SEP 0 52000

BHT-412-FMS-45.3 AND 45.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODELS ^S12/S^1ZLL

ROTOR CRAFT

FLIGHT MANUAL

SUPPLEMENT IN> D MAXIMUM CONTINUOUS IN6 POVER IKWTH PT6T-3D ENGINE 706-029) "V/,~/? BHt3' BHT 36

THUH

36080 UENT

,yON TEMPORA"F MAST TORQUE SIG NpNDIONER ASB 412This Tempoary Revison etends perfmnan and d provides new cmpTace with ASB 412-97-0.

nse these teemporay revision pages oppote II pages in Right Manual after compliance with ASB 41 DO NOT remove existing pages from Right Manual.

Inlormation contained herein supplements informaton of basic Flight Manual. For Limitations, Procedures, and Perfomance Data not contained in this supplemnt, consult basic Right Manual.


Bell Helicopterptir il

COPl T NOTICE

TEXTCM IA Dl OIf

O ELLI IELCOPTER TEM

A Suidilry ot Ttxron I N.

L

23 JUNE 1994

ALLRITS RESrVED


T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-45.3 & 45.4

LOG OF REVISIONS

Original...........0 .............. 27 Aug 93 Reissue ........... 0.............. 23 Jun 94 Revision..........1 ............... 26 Apr 95

Revision ......... 2 ............. 08 AUG 97 Revision .......... 3 ............. 05 SEP 00

LOG OF PAGES

REVISION NO.

REVISION

PAGE

PAGE

.NO.

T itle ..................................... N P ...................................... A - B .................................. 1 ........................................

3 0 3 1

2 ....................................... 3 3/4 ...................................... 3 5- 26 .................................. 3

NOTE


SEP 052000

Rev.3

A

T N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-45.3 AND 45A



NOTICE

This Temporary Revision has been cancelled. Subject data has been incorporated in BHT-412-FMS-45.3 and 45.4 at Revision 3. Remove and discard all Temporary pages dated 8 August 1997 from this manual including these canceled pages.

APPROVED:

MANAGER

/

SEP

5 2000

ROTORCRAIT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

TEMPORARY REVISION - 8 AUGUST 1997 B CANCELED - 5 SEP 2000



Original........... 0 .............. 27 Aug 93 Reissue........... 0.............. 23 Jun 94 Revision..........1 ............... 26 Apr 95

DATE:

APPROVED:

MANAGER


B

Rev. 3

08 AUG 97 Revision .......... 2 ............ Revision .......... 3 ............. 05 SEP 00

SEP 05 2000

T N I 1 R Section P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-45.3 & 45.4

FAA APPROVED

1-1. INTRODUCTION

CAUTION

Limitations not listed herein are same as those shown in basic flight manual, BHT412-FM-3 or BHT-412-FM-4.

1-5. CONFIGURATION

1-5-B. OPTIONALEQUIPMENT

WHEN OPERATING NEAR THE MAXIMUM MAST TORQUE LIMIT, OVERTORQUE INADVERTANT DURING MAY OCCUR FLIGHT MANEUVERING CONDITIONS INVOLVING TURNS AND/OR NOSE DOWN ATTITUDE CHANGES. DECREASE POWER TO 90% MAST TORQUE PRIOR TO MANEUVERING HELICOPTER.


Helicopters S/N 36020 through 36080 must be equipped with PT6T-3D engines. Refer to BHT-412-FMS-46.3 for additional engine limitations.

1-14. TRANSMISSION 1-14-C. MASTTORQUE 1-14-C-1.

TWIN ENGINE OPERATION

Maximum with airspeed up to 105 KIAS

Maximum with airspeed above 105 KIAS

100%

81%

Rev.

1

I

T N I R P N O R I O T F C U T O OD N R E P E E R R C S Section 2 N G O N I R W O E I F Y V L N 0 Section 3 O FAA APPROVED

BHT-412-FMS-45.3 & 45.4

TRIPLE TORQUE

MAST (TWIN ENGINE OPERATION) Continuous

10 81 to

81 to 100%

5 to

Figure 1-1.

73.2%

2 1/2 minute OEI

81

Maximum OEI

Instrument markings


Rev.3 Rev. 3

SEP 0 52000

Continuous OEIoperation

73.2 to 81%


2

Continuous for airspeed

T N I Section 4 R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

4-1.

INTRODUCTION

Performance data presented herein are derived from engine manufacturer's

specification power for PT6T-3D engine less installation losses.

4-14. PERFORMANCE CHARTS

Increased twin and single engine performance can be achieved as presented herein. Performance charts that differ from those in BHT-412-FMS-46.3 or BHT-412-FM4 are as follows: Hover ceiling, OGE, at maximum continuous power (figure 4-2)

BHT-412-FMS-45.3 & 45.4

Single engine rate of climb at maximum continuous OEI power (figure 4-4)

Hover ceiling and rate of climb charts at

takeoff power and 30-minute OEI power in

basic manual should be disregarded since those power ratings do not apply to this helicopter configuration.

NOTE

Published single engine performance is intended for emergency use only when one engine becomes inoperative due to an actual malfunction. Routine operation in an OEI range can affect engine warranty and service life.

Twin engine rate of climb at maximum continuous power (figure 4-3)

SEP 052000

Rev.

3

3/4



BHT-412-FMS-45.3 & 45.4


SKID HEIGHT 60 FT. HEATER OFF ° -40 TO 0 C

MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR 150 AMPS(EA.) CAUTION

-40

-30

OGE HOVER OPERATION MAY RESULT IN VIOLATION OF H-V LIMITATIONS

-10 -20 OAT - °C

0

11 12 10 9 8 GROSS WEIGHT - LBS x 1000

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

412FS-45 34 2-2

Figure 4-2.

6

Rev. 3

SEP 05 2000

Hover ceiling (Sheet 2)


BHT-412-FMS-45.3 & 45.4



MAXIMUM CONTINUOUS POWER ENGINE RPM 100% GENERATOR 150 AMPS(EA.) CAUTION

OGE HOVER OPERATION MAY RESULT IN VIOLATION OF H-V LIMITATIONS

14,000 FT HD LIMIT

0 10

20

8

9

11

10

1

Figure 4-2. Hover ceiling (Sheet 3)

SEP 052000 Rev.

3

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-45.3 & 45.4

FAA APPROVED

HOVER CEILING





3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

412FS 45.3-4-2-4

Figure 4-2.

8

Rev. 3

SEP 052000

Hover ceiling (Sheet 4)


BHT-412-FMS-45.3 & 45.4


70 KIAS HEATER OFF





0

1 2

3

4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

20000

14000

6000

4000

2000

400

800 1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FT/MIN

412FS-45.3-4-3-1

Figure 4-3.


Rev. 3

9


FAA APPROVED



70 KIAS


HEATER ON



RATE OF CLIMB -

0 1 2

3

4 5 6

M/SEC

7 8 9 10 11 12 13 14 15 16 17 18

20000 18000

16000

14000 14000

6000

4

2000 0

412FS-45.3-4-2

Figure 4-3,

10

Rev. 3


SEP 05 2000


BHT-412-FMS-45.3 & 45.4



70 KIAS HEATER OFF

WTH ALL DOORS OPEN OR REMOVED:

0

0

1 2 3 4

400

800



1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB

-

FT/MIN

412FS-45.3-4-3-3


SEP 052000

Rev.3

11


FAA APPROVED



70 KIAS HEATER ON



RATE OF CLIMB

0

1 2 3

4 5

-

M/SEC

6 7 8 9 10 11 12 13 14 15 16 17 18

20000 18000

16000

12000

12000

1000 8000 6000

412FS-45.3-4-3-4

4000

2000

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

Figure 4-3.

12

Rev. 3

FT/MIN


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L o N * O BHT-412-FMS-45.3 & 45.4

FAA APPROVED


MAX COTNUU

!

70 KIAS HEATER OFF




012RATE

B

0

3

1 2

4 5 6

OF CLIMB - M/SEC 7 8 9 1011 12131415161718

20000

6

18000

16000 1 -

6

I,

-- x~,'x I\ ~

14000

j 12000

-

8 _3 9

_

_

4

X

)

-

_ U^

2|\_

a_ 1000 D

t"L48000 _ 6800000

10

-10

20

20

20

-

-_

Ie3 S2000P 052000

0

400

800

13


412 F$-43-4-3-5

Figure 4-3.


SEP 052000

Rev. 3

13


FAA APPROVED



70 KIAS

ENGINE RPM 100% GENERATOR 150 AMPS (EA)

HEATER ON


1. CLIMB SPEED IS 60 KIAS 2. RATE OF CUMB WILL DECREASE 275 FT/MIN

RATE OF CLIMB

-

M/SEC

0 1 2 3 4 5 6 7 8 9 10 1112 13 14 15 16 17 18

20000

18000

16000

14000

12000

10000

3

L

8000

4000

0

400

800

1200 1600 2000 2400 2800 3200

RATE OF CLIMB

-

FT/MIN

412FS-45.3-4-3-6

Figure 4-3. 14

Rev. 3

SEP 05 20 0 0


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y . 2000 V L N O FAA APPROVED

BHT-412-FMS-45.3 AND 45.4



70 KIAS HEATER OFF


0

1 2

4

3


RATE OF CLIMB - M/SEC 5 6 7 8 9 1011 12131415161718

20000

18000

16000

14000

12000

10000

6000 6000 6000

-2

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB

-

FT/MIN

412FS-45.3-4-3-7 Figure 4-3.

SEP 05 2000

Twin engine rate of climb (Sheet 7) Rev. 3

15


BHT-412-FMS-45.3 AND 45.4




70 KIAS

HEATER ON


1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMS WILL DECREASE 275 FT/MIN


RATE

0

16

Rev. 3

1 2

5

4

5

6

OF CLIMB

7

8

-

M/SEC

9 1011 12131415161718

SEP 05 2000


BHT-412-FMS-45.3 AND 45.4



70 KIAS

ENGINE RPM 100%

HEATER OFF



0

1 2 3

0

400



800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

FT/MIN

412FS-45.3-4-3-9

Figure 4-3.

SEP 052000


Rev.3

17


FAA APPROVED




70 KIAS HEATER ON


0

1 2

3

4


RATE OF CLIMB - M/SEC 5 6 7 8 9 1011 1213 1415161718

20000

18000

12000

8000 6000

4000

2000

0

0

400

800

1200 1600 2000 2400 2800 3200 3600

RATE OF CLIMB -

FT/MIN

412FS-45.3-4-13-10


18

Rev. 3

SEP 052000


FAA APPROVED



70 KIAS HEATER OFF




RATE OF CLIMB -

0

M/SEC

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

20000

18000 16000

14000

2

6000

I

412FS-45.3-4-3-11


SEP 052000

Rev. 3

19


FAA APPROVED




70 KIAS HEATER ON


1. CLIMB SPEED IS 60 KIAS 2. RATE OF CUMB WILL DECREASE 275 FT/MIN

RATE OF CLIMB -

0 1 2 3

4

5 6

M/SEC

7 8 9 1011 12131415161718

2000 18000

16000

14000

0

12000

10000

6000

RATE OF CLIMB -

Figure 4-3.

20

Rev. 3

FT/MIN


SEP 0 5 2000


FAA APPROVED

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 7000 LB (3175 kg)


70 KIAS





RATE OF CLIMB

-4-3-2-1

0

1

2

3

4

-

5

M/SEC

6

7

8

9 10 11 12

20000

18000

16000 14000

10000

8000

2

6000

4000 2000

-800

-400

0

400

800

RATE OF CLIMB

Figure 4-4.

1200 1600 - FT/MIN

2000

2400

Single engine rate of climb (Sheet 1 of 6)

SEP 05 2000

Rev. 3

21


BHT-412-FMS-45.3 & 45.4






RATE OF CLIMB -

RATE OF CLIMB

Figure 4-4.

22

Rev. 3

M/SEC

-

FT/MIN


SEP 05 2000


BHT-412-FMS-45.3 & 45.4


70 KIAS






RATE OF CLIMB -

-1600-1200

-800

RATE -400OF CLIMB 0

M/SEC

FT/MIN - 400 800

1200

1600


SEP 05 2000

Rev. 3

23


FAA APPROVED

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB (4536 kg)





RATE OF CLIMB

-10-9 -8 -7 20000

000

-

M/SEC

-6 -5-4 -3-2-1

0

1

2

3

4

5

6

TWIN ENGINE MCP

6000

14000

a

6000

4000 2000

- 2000 -1600-1200 -800

-400

RATE OF CLIMB Figure 4-4.

24

Rev. 3

400

-

FT/MIN


800

1200


BHT-412-FMS-45.3 & 45.4






RATE OF CLIMB - M/SEC -10-9-8-7-6-5-4-3-2-1 0 1 2

-2000 -1600

-1200

Figure 4-4.

3

-800 -400 0 400 RATE OF CLIMB - FT/MIN

4

800

5

6

1200


SEP 05 2000

Rev. 3

25


FAA APPROVED






RATE OF CLIMB

-

M/SEC


Figure 4-4.

26

Rev.3


SEP 052000

BHT-412-FMS-46.3

T N I R P N ROTORCRAFT O I *t OR FLIGHT MANUAL T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O *y

MODEL ^1

SUPPLEMENT

PT6T-3D ENGINES 412-706-049

BHT 36020 THROUGH 36086 CERTIFIED 14 DECEMBER 1993

This supplement shall be attached to Model 412 Flight Manual when PT6T-3D engines have been installed.


CPITI COPYRIGHT NOTICE 2000 COPYRIGHT INC. TEXTRON BELL® HELICOPTER ANDBELLHELICOPTER TEXTRON INC. A DIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

Bell Helicopter H ii

t

Inc A Sub.diary ofTextron

POSOFFICEBOX482· FORTWORTH.TEXAS76101

14 DECEMBER 1993 REVISION 9-5 SEPTEMBER 2000


NOTICEPAGE


T N I MODEL ^Sf R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

ROTOR CRAFT FLIGHT MANUAL SUPPLEMENT PT6T-3D ENGINES

(412-706-049)

20 THROUGH 36086

REVISION L CONDITIONER

ETEp MAST TORCGI E/

This Temporary Revbio performance and provides compliance with ASB 412-97-

fe.,cent Yioitolr

k

published data after

like numbered go Attach these temporary revision ASB41 pages in Flight Manual after complia DO NOT remove existing pages from Flig

Information contained herein supplements ino Flight Manual. For Limitations, Procedures, and Data not contained in this supplement, consult Manual.

n


Bell Helicopteru

lJ-]

A Substi.ar of Tetron It oniornC hOX'IB

NOTICE CDWlHGTr

BU lKICR[ TEXTRGNlNC. D

TENs C

LTD

ALL RIQOTS RESERVED

®

FOIrWOItDECEMBER

1

14t DECEMBER 1993 TEMPORARY REVISION - 8 AUGUST 1997 CANCELED - 5 SEPTEMBER 2000


FAA APPROVED


Temporary ..................... Canceled .......................

8 AUG 97 5 SEP 00

LOG OF PAGES

Revision No.

Page

Title ........................ A- B ...................... 10-18 .................... 18A -18B .................. 18C/18D .................... 19- 48 ....................

Revision No.

Page

Temporary Temporary Temporary Temporary Temporary Temporary

NOTICE

This Temporary Revision has been canceled. Subject data has been incorporated in BHT-412-FMS-46.3 at Revision 9. Remove and discard all Temporary pages dated 8 August 1997 from this supplement including these canceled pages.

APPROVED:

MANAGER

ROTORCRART CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

A


DATE:

SEP

52000

T N I OD R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

LOG OF REVISIONS

Original ........... 0 .............. 14 Revision ................. 14 Revision .......... 2 .............. 23 Revision .......... 3 ............. 29 Revision .......... 4 .............. 13

DEC FEB JUN AUG NOV

93 94 95 95 95

Revision..........5 Revision..........6 Revision .......... 7 Revision .......... 8 Revision..........9

.............. ............. .............. ............. .............

23 FEB 30 AUG 13 JAN 08 AUG 05 SEP

96 96 97 97 00

LOG OF PAGES

*PAGEb~ ~REVISION

PAGE

NO.

Title ..................................... NP ...................................... AB .................................. i/ii ....................................... 1-2 ................................... 2A12B ................................... 3-4 ................................... 5 .......................................

*^&b~~~

9 0 9 2 4 4 4 0

REVISION NO.

PAGE

6 ....................................... 6A - 6B ............................... 6C/D .................................... 7-9 ................................... 10- 18 ................................ 18A - 18B ............................. 18C/D ................................... 19- 48 .................................

9 9 9 0 9 9 9 9

~NOTE


SEP 05 2000

Rev. 9

A



Original........... 0..............14 DEC Revision .......... 1 .............. 14 FEB Revision..........2 .............. 23 JUN Revision .......... 3 ............. 29 AUG Revision .......... 4 .............. 13 NOV

APPROVED:

MANAGER

Rev. 9

Revision Revision Revision Revision Revision

.......... 5 .......... 6 .......... 7 .......... 8 .......... 9

DATE:

SEP

.............. 23 FEB ......... . 30 AUG .............. 13 JAN ............. 08 AUG ............. 05 SEP

96 96 97 97 00

5 2000

/


B

93 94 95 95 95

rSrp52

t


BHT-412-FMS-46.3

GENERAL INFORMATION

The PT6T-3D Engines enable the helicopter to achieve Increased OEI performance.

Rev.2

i/ii


BHT-412-FMS-46.3

INTRODUCTION

The PT6T-3Dengines permit increased OEI performance. Refer to Figure 1-1. Refer to BHT-412-FM-3for limitations not provided In this supplement.

WEIGHT/CG

logbook detailing the extent of operation in excess of twin engine limits. This does not apply to approved ITT limits for starting.

GAS PRODUCERRPM(N1 )

TWINENGINEOPERATION

Actual weight change shall be determined after engine is installed and ballast adjusted, If necessary, to return empty weight CG to within allowable limits.

AMBIENTAIR TEMPERATURE

Minimum ambient temperature for operation at all altitudes with engine oil pressure/temperature Indicator 209-070262-113Installed is -40°C (-40°F). Minimum ambient temperature for operation at all altitudes with engine oil pressure/temperature indicator 209-070262-109installed is -34°C (-30°F). NOTE Refer to Oil Pressure Limits.

POWERPLANT

NOTE

Operation in an OEI power range is Intended for emergency use only when one engine becomes due to an actual inoperative malfunction

Anytime an engine is operated In an OEI range, an entry shall be made in helicopter

Continuous Maximum Continuous

61 to 103.1% 103.1%

ONE ENGINE INOPERATIVE (OEI) Continuous OEI 21/2-minuteOEI Maximum OEI

103.1to 103.7% 103.7to 109.2% 109.2%

INTERTURBINE TEMPERATURE (ITT)

TWINENGINEOPERATION

GAGE PN 212-075-067-115 Continuous 300 to 810 °C Maximum continuous 810 °C Maximum transient 925 °C Do not exceed 5 seconds above 810 °C. Maximum for starting. 1090 °C Do not exceed 2 seconds above 960° C. Rev.4

1


BHT-412-FMS-46.3

NOTE Either ITT gage may be Installed.

GAGE PN 212-075-067-119 300 to 810 °C Continuous 810 °C Maximum continuous 940 °C Maximum transient Do not exceed 5 seconds above 810 °C. Maximum for starting. 1090°C Do not exceed 2 above seconds 960° C.

GAGE PN 209-070-262-113 40 psi Minimum for idle 40 to 90 psi Operation below 79% N1 RPM 90 to 115 psi Continuous operation 115 psi Maximum GAGE PN 209-070-262-109 40 psi Minimum for idle 40 to 80 psi Operation below 79% N1RPM 80 to 115 psi Continuous operation 115 psi Maximum

OILTEMPERATURE

CAUTION

INTENTIONALUSE OF ITT ABOVE 810 °C IS PROHIBITED DURING NORMAL OPERATIONS EXCEPT DURINGSTART.

ONEENGINEINOPERATIVE

GAGE PN 212-075-067-115 820 °C Maximum continuous OEI 820 to 925 °C 2 1/2Minute OEI Maximum OEI 925 °C NOTE Either ITT gage may be installed.

GAGE PN 212-075-067-119 820 °C Maximum continuous OEI 820 to 940 °C 21/2Minute OEI 940 °C Maximum OEI

OIL PRESSURE ENGINE

NOTE Either Engine Oil Pressure/ gage may be Temperature installed. Refer to Ambient Air Temperature limits.

2

Rev.4

GEARBOX ENGINEANDCOMBINING Minimum Continuous Maximum for MIL-L7808 oil Maximum for MIL-L23699 or DOD-L85734 oil

0 °C

0 to 115°C 115°C 120°C

ENGINETORQUE

ONEENGINEINOPERATIVE Continuous Maximum Continuous 2/2 Minute Maximum

5 to 73.2% 73.2 % 73.2 to 81% 81%

FUEL AND OIL

OIL GEARBOX ENGINEANDCOMBINING Oil conforming to PWA Specification 521 Type I and MIL-L-7808 (NATO 0-148) limited to ambient temperatures from -40°C (-40°F) to 43°C (110°F). Oil conforming to PWA Specification 521 Type II and MIL-L-23699 (NATO 0-156) limited to ambient temperatures from -40°C (-40°F)to 51.7°C (125°F).

PLACARDSAND DECALS Refer to figure 1-2.

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O A

12%

T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O


BHT-412-FMS-46.3

Figure 1-l.

Instrument

markings

(Sheet 3 of 3)


FAA APPROVED

TWIN 2.5 MIN OEI

103.1% 109.2%

CONTINUOUS

OEI

103.7%

412-FMS-46.3-1-2

Figure 1-2. Placards

5

T N I e Section 2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-46.3

ENGINE START

2-5.

2-5-A.

MOISTURE IN THE ENGINE FUEL CONTROL SYSTEM. MONITOR ENG RPM (N,,) DURING COLD WEATHER STARTING FOR

ENGINE START

OVERSPEED. IF AN OVERSPEED

APPEARS IMMINENT, ABORT START AND CLOSE THROTTLE TO OFF POSITION.

NOTE

Either engine may be started first. However, the following procedure is provided for starting engine 1 first.

START switch - ENG position. Observe starter limitations.

At ambient temperatures 5 °C (41 °F) and below, refer to paragraph 2-5-F.

Engine 1 ENGINE Indicating.

Engine 1 FUEL TRANS switch - ON, check NO. 1 FUEL TRANS caution light extinguished.

Engine throttle -

Engine 1 BOOST PUMP switch - ON, check NO. 1 FUEL BOOST light extinguished.

Engine 1 ITT - Monitor to avoid a hot start. Maximum ITT during start is 1090°C, not to exceed two seconds above 960°C. If ITT continues to rise, abort start by activating idle stop release and rolling

throttle to fully closed. Starter should remain engaged until ITT decreases. Do not attempt restart until corrective maintenance has been accomplished.

caution light will illuminate momentarily.)

Rotor -

Check.

Clear.

NOTE

---------CAUTION _****«»**««***,

PROLONGED EXPOSURE TO AMBIENT TEMPERATURES OF O °C (32 °F) OR LESS MAY FREEZE

6

Rev. 9

Open to idle at 12% GAS

PRODUCER RPM (N|) minimum.

switch 1 FUELN. ( E Engine switch - ON. (FUELL VVALVE Engine 1 FUEL

Engine 1 FUEL PRESS indicator -

OIL pressure -

SEP 05 2000

_

If engine fails to start, refer to paragraph 2-5-D procedures of Basic Flight Manual.


Ensure in full down

T N I D R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L 0w N O FAA APPROVED

BHT-412-FMS-46.3

POSITION. VERIFY PROPER OPERATION PRIOR TO FLIGHT.

CAUTION

IF STICK CENTERING INDICATOR

Engine 1 throttle - Increase to 77 - 85% ENG RPM (N,,). Friction as desired.

GROUND OPERATION SHALL BE CONDUCTED AT 97% ROTOR RPM OR ABOVE.

NOTE

SYSTEM

IS

INOPERDATIVR

Minimize blade flapping by maintaining highest rotor RPM (NR) within allowable range.

NOTE

On side slopes greater than five degrees, disregard CYC CTR caution lights and position cyclic as desired.

Rotor RPM (NR) desired.

Cyclic - Position as necessary to extinguish CYC CTR caution lights. NOTE

~~~~NOTE ~used,

GAS PROD - Check 61 ± 1% when throttle is on idle stop. NOTE

t~C~............t

DURING RPM INCREASE, ANY ABNORMAL INCREASE IN ONEPER-REV VIBRATION MAY INDICATE ONE OR MORE MAIN DROOP RESTRAINTERS FAILED TO DISENGAGE FROM STATIC

NOTE Before attempting generator assisted start on second engine, it is recommended that the battery be charged until the ammeter load

AMPS 1 indicator amps.

OIL

Engine 1 PART SEP OFF caution light Check extinguished. :U,;

On.

drops below 150 amps.

Do not increase ROTOR above 85% RPM until XMSN oil temperature is above 15 °C.

CAUTION

Check 71% minimum.

GEN 1 switch-

START switch - Centered (off) at 55% GAS PROD RPM (N,).

ENGINE, XMSN, and GEARBOX PRESSURES - Check.

If external power is used, proceed to engine 2 start. If battery was Proceed as follows:

N| RPM -


Minimum 77 - 85%, as

2-5-B.

Check at or below 150

ENGINE 2 START

NOTE NOTE At ambient temperatures of 5 °C (41 °F) and below, refer to paragraph 2-5-F.

Engine 2 FUEL TRANS switch - ON. Check No. 2 FUEL TRANS caution light extinguished.

Engine 2 BOOST PUMP switch - ON. Check No. 2 FUEL BOOST light out (FUEL XFEED caution light will illuminate monentarily.

SEP 05 2000

Rev. 9

6A

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W 0 O E I F Y V L N 0 O FAA APPROVED

BHT-412-FMS-46.3

Section 3

K^^&


0

SEP 0 5 20X

Rev. 9

6C/D

T N I Section4 PR N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-46.3

INTRODUCTION

Performance data presented herein are derived from engine manufacturer's specification power for PT6T-3D engine less installation losses.

emergency use only when one engine becomes inoperative due to an actual malfunction. Routine operation In an OEI range can affect engine warranty and service life.

HOVERCEILINGOGE

POWERASSURANCECHECK

Power Assurance Check charts for engine are presented in figure 4-1.

The hover ceiling charts (Figure 4-2) provide maximum weights for hovering OGE at all pressure and density and outside air temperature conditions with heater on or off.

CAUTION

Do not exceed maximum continuous limits (103.1% N,) RPM, 810°C ITT, or 73.2% torque) during power assurance checks.

PERFORMANCE CHARTS

Increased twin and single engine performance can be achieved as presented herein. Performance charts that differ from those in BHT-412-FM-3 are as follows:

TWIN ENGINE CLIMB

RATE

OF

Twin engine climb rates for maximum continuous engine power, maximum continuous transmission power, and takeoff transmission power ratings are shown In (Figure 4-3).

SINGLE ENGINE RATE OF CLIMB

Single engine climb rates are computed on single engine maximum continuous power rating (Figure 4-4).

NOTE

Published performance

single engine is intended for

7


FAA APPROVED

MODEL412 POWERASSURANCE CHECK- HOVER PT6T-3D ENGINE

HEATER/ECU- OFF.

IF INDICATED ITT EXCEEDS MAX ALLOWABLE, REPEAT CHECKSTABILIZINGPOWERFOURMINUTES.

THROTTLES: TESTENGINE- FULL OPEN, FRICTIONED. OTHERENGINE- FUGHT IDLE.


Nl RPM - 97%.

COLLECTIVEPITCH - INCREASEUNTIL LIGHT ON SKIDS OR HOVERING.DO NOT EXCEED810' ITT, 103.1% NI RPM, OR 73.2% TORQUE.

IF EITHER ENGINE EXCEEDS ALLOWABLE ITT AFTER STABILIZING FOUR MINUTES, PUBLISHED PERFORMANCE MAY NOT BEACHIEVABLE.CAUSESHOULDBE DETERMINED AS SOON AS PRACTICAL

STABILIZEPOWER ONE MINUTE, THEN RECORDPRESSURE ALTITUDE,OAT,ENGINETORQUE AND ITT. ENTERCHART AT INDICATEDENGINETORQUE,MOVE UP TO INTERSECTPRESSUREALTITUDE. PROCEEDTO THE RIGHT TO INTERSECTOUTSIDEAIR TEMPERATURE, THENMOVE UP TO READ VALUEFOR MAXIMUM ALLOWABLEITT.

MAXIMUM ALLOWABLEITT - °C 600 650, 700 750


412-FMS-46.3-4-1-1

Figure 4-1.

8

Power assurance check (Sheet 1 of 2).


FAA APPROVED

MODEL 412 POWERASSURANCE CHECK - IN-FUGHT PT6T-3DENGINE

ESTABLISH LEVELFLIGHTABOVE 1000 FEETAGL AIRSPEED- 100 KIAS (ORVNE, IF LESS).

ENTERCHARTAT INDICATEDENGINETORQUE,MOVE UP TO ALTITUDE, PROCEEDTO THE RIGHT INTERSECTPRESSURE THENMOVE UP OUTSIDEAIRTEMPERATURE, TO INTERSECT TO READVALUEFORMAXIMUM ALLOWABLEITT.

HEATER/ECU- OFF.


THROTTLES: TEST ENGINE- FULL OPEN,FRICTIONED.


OTHERENGINE- DECREASESLOWLYUNTIL TESTENGINE TORQUEIS WITHIN TESTRANGE.DO NOT EXCEED810°C ITT, 103.1% NI RPM,OR 73.2% TORQUE.

NII RPM - 97%.

IF EITHER ENGINE EXCEEDS ALLOWABLE ITT AFTER STABILIZINGFOUR MINUTES, PUBLISHEDPERFORMANCE MAYNOT BEACHIEVABLE.CAUSESHOULDBEDETERMINED ASSOON AS PRACTICAL.

STABILIZEPOWER ONE MINUTE IN LEVEL FLIGHT, THEN RECORDPRESSUREALTITUDE,OAT,ENGINETORQUE AND ITT.

500

40 50 TORQUE ENGINE

550

MAXIMUM 600

80 (INDICATED) - 60 PERCENT 70

Figure 4-1.

ALLOWABLE ITT - °C 750 700 650

800

412-FMS-46.3-4-1-2

MAXIMUMCONTIUOS

412-FMS-40.3-4-1-2


9


FAA APPROVED





14,000 FT HD LIMIT

O

10

20

30

OAT - °C

40

50

I I I I I IIII

8

9

10

GROSS WEIGHT -

3.4

3.8

4.2

10

Rev. 9

SEP 0 5 2000


I___ -I

12

LBS x 1000

4.6

GROSS WEIGHT -

Figure 4-2.

11

5.0

5.4

KG X 1000

v


BHT-412-FMS-46.3



SKID HEIGHT 60 FT. HEATER OFF -40 TO 0° C


14,000 FT HD LIMIT

*|

~

-20

, -

/~~

-30

-20 OAT w~x'~I

~~~~

~

_A

1,/

I>Ii

. 38424.

,~,

~ ~~GOSWIH

-10 C

0

0

.

GX10

.

8 9 10 11 12 GROSS WEIGHT - LBS x 1000 I I I I I I'7 3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT -

Figure 4-2.

C

9 'SEP 05 2000 10 ARe. WIHT-,/

CGRS

*OT-

.,./

J

AR A-A

-40

OT -

_S_

-S/~^

KG X 1000


1

T N I R P N O R I O T F C U T O OD N R E P E E R R C _--_i-ijjj S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

FAA APPROVED

HOVER CEILING


TAKEOFF POWER SKID HEIGHT 60 FT. ENGINE RPM 100% HEATER ON GENERATOR 150 AMPS(EA.) 0 TO 20°C CAUTION: OGE HOVER OPERATION MAY RESULT INVIOLATION OF H-V LIMITATIONS

14.000 FT HD LIMIT

-

AREAR A

7_

|

_

0

10 20 8 9 10 11 12 OAT - °C GROSS WEIGHT - LBS x 1000 I

3.4

I I

3.8

I

4.2

I

4.6

GROSS WEIGHT -

Figure 4-2.

12

Rev. 9

SEP 0 5 2000

I

I

5.0

'

I

5.4

KG X 1000



BHT-412-FMS-46.3

HOVER CEILING



SKID HEIGHT 60 FT. HEATER ON -40 TO 0° C


14,000 FT HD LIMIT

AREAA

b

/

-40

C

/-

-30

-20

OAT -

.

,

_

-10

0

°C

/ __

8

___ -

9

10

Ir

3.4

_

3.8

I'_

4.2

I

LBS x 1000

4.6

GROSS WEIGHT -

12

11

GROSS WEIGHT -

5.0

I

5.4

KG X 1000

Figure 4-2. Hover ceiling (Sheet 4 of 12)

,SEP

0520M

Rev.9

13

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY 0 O BHT-412-FMS-46.3

FAA APPROVED

HOVER CEILING





14,000 FT HD LIMIT

AREA A -

0

10

20

OAT - °C

8

9

10

11

12

GROSS WEIGHT - LBS x 1000 I mI I I I I

3.4

3.8

4.2

4.6

5.0

5.4


Figure 4-2.

14

Rev. 9

SEP 05 2000


0

T N I R P N O R I O T F C t U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-46.3

HOVER CEILING



SKID HEIGHT 60 FT. WINTERIZATION HEATER ON -40 TO 0° C


14,000 FT Ho UMIT

,~';/

-

/

~

b.

-40

-43

~3V

~

,.

3.8

3.4

*

AT-4C,

~

4.2

4.6

5.0 0

9

10

11

5.4

/

-30

-20

-10

OAT -ÔAT °C

0

-C

8

3.4

3.8

4.2

4.6

GROSS WEIGHT -

Figure 4-2.

12

LBS U3S xx 1000 1000

GROSS GROSS WEIGHT -

5.0

5.4

KG X 1000


;SEP 0 520%0

Rev.9

15


FAA APPROVED

HOVER CEILING





14,000 FT Ho LIMIT

1020 ,, 2 ~.~y89

. ~ ,.;.."-0

__40

_

0

I

16

0

10

20

2000

940

50

'

8

I '

I

' I

10

I

11

I

I

12


BHT-412-FMS-46.3



SKID HEIGHT 60 FT. HEATER OFF -40 TO O°C


OAT - °C

14.000 FT Ho LIMIT

-/__

_

--

- --__

*--T^-^----3,-20 1_ __ / ,

/_ /

-40

-30

-20

-10

OAT - °C

-e9 10

*A-S-P 0 8 --- A 9 I

I

I I

11

12

LBS x 1000

GROSS WEIGHT -

I

[

I

]

I

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

JJ^^^~~~

~

Figure 4-2. Hover ceiling Figure 4-2. Hover ceiling (Sheet (Sheet 88 of of 12) 12)

SEP 0520

Rev. 9

17


FAA APPROVED

HOVER CEILING





14,000 FT HD UMIT

/

0

AREA

A

10 20 8 9 10 11 12 OAT - °C GROSS WEIGHT - LBS x 1000 I

Ii

'

I

'

I

'I

3.4 3.8 4.2 4.6 5.0 5.4 GROSS WEIGHT - KG X 1000

Figure 4-2.

18

Rev. 9

SEP 05 ZDD0


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L 0w~~~~N O FAA APPROVED

BHT-412-FMS-46.3

HOVER CEILING





OAT - °C

14,000 FT HD LIMIT

_

*-s.^

-40

-30

-20

OAT -

-10

C

0

_~_ rv~b /_ _ _ _ _ _30O

5^

8

9

10

11

GROSS WEIGHT -

I ' I' I 3.4 3.8 4.2

I

I 4.6

GROSS WEIGHT -

SEP 05200O 052 2a1 S£P

12

LBS x 1000

I 5.0

7

I 5.4

KG X 1000

Rev.9 Re..9

18A


FAA APPROVED





14,000 FT HD LIMIT

--r

0

-I_I -1|

-I |-

10

20

OAT - °C

II ~

8

9

10

GROSS WEIGHT -

I II I 3.4 3.8

4.2

1---

tn

I

I

4.6

11

_

12

LBS x 1000 I

I

5.0

5.4


Figure 4-2.

18B

Rev. 9

SEP 0 5 2000


0


BHT-412-FMS-46.3





14,000 FT HD LIMIT

OAT -

C

/_>^_

OATZ-

GROSS WEIGHT - LOS40 x 100

I ' ' I '- IKG' XI 1000 WEIGHT /-GROSS ____AREA A

GOSI_ EGH_TI_ :

GROSS WEIGHT -

Figure 4-2.

x

KG X 1000


SEP 0 52 D2

Rev. 9

18C14-18D


BHT-412-FMS-46.3


TAKEOFF POWER

70 KIAS


HEATER OFF


~)^^b~

0

~RATE

1 2

20000 III ""11 20000"

3 4 5


OF CLIMB - M/SEC 6 7 8 9 1011 12 131415161718 ll

ll ,iIimHlilml

llilmlmI

ll 'lnlhlml ,mnmil

18000

16000

m

IoooI-

----

---- ^\

S

14000

I

,J 12000

10000

we

8000

000 \0000

4-.

RATE OF CLIMB

Tn e

e re of

-

30

FT/MIN

STEP O52D 24)

3

Rev.-2

<

19N

Figure 4-3. Twinengine rate of climb (Sheet I of 24)

,SEP 0 5 200

Rev. 9

19


FAA APPROVED



70 KIAS HEATER ON



RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 1011 12 1314 15161718 u l'i Ji' n illl

0

1 2 U 20000 "

18000

-

16000

1'

14000

2,

12000 9

'

6000

12

-

\

\

x

',,,'-,

_

_

_

--

_

_

_

_

_

_

_

_

_

_5_

:

1

4000

/!

2000

0

400

800

1200

1600 2000 2400 2800 3200 3600

RATE OF CLIMB

Figure 4-3.

20

Rev. 9

-

FT/MIN


SEP 05,'Qe


BHT-412-FMS-46.3


TAKEOFF POWER

70 KIAS


HEATER OFF


~RATE

JJ^yt~

0

3 4 5 6

1 2

m

20000

~V


WU

\Jl

OF CLIMB - M/SEC 7 8 9 1011 12 1314 15161718 U

l

i

u

ll

l

i,,il6

~18000

__

16000

5(\,__

w 14000

L

0520

_EP

1000

Rev.9

210

--

'\ \

"'"8000\ LJ

- 5<

'<'/l

I

\

4000

2000

'-

0

400

l

800

1200

1600 2000 2400 2800

RATE OF CLIMB

Figure 4-3.

-

3200

3600

Fr/MIN


'SEP 0 52009

Rev. 9

21


FAA APPROVED



70 KIAS HEATER ON



RATE OF CLIMB -

0

1 2 3 4 5 6 7

20000

18000

M/SEC

8 9 10 11 12 13 14 15 16 17 18

ULt Wil'l.'ILj

i

S

-

0

14000

Il

12000 2"

_2000I_

3a .-

10000<

22 22

Rev.69 Rev. 9

SEP 0 5 2o0 SEP052008


BHT-412-FMS-46.3


70 KIAS

TAKEOFF POWER


HEATER OFF


0

1 2

20000

18000

0b

3

1. CUMB SPEED IS 60 KIAS

2. RATE OF CLIMB WILL DECREASE 275 FT/MIN

RATE OF CLIMB - M/SEC 5 6 7 8 9 10 11 12 13 14 15 16 17 18

4

'l

i

'

I

'l

i

6tt

-

.

16000 _

\"

" 16000 _ _ dj 14000 ---

LjL

IO

___s-N ^

._

1[ 1 I i

__

----

000 0

12000

U)

10000

400

800

1200

1600 2000 2400 2800 CLIMBd F-/M -

OF

200

60 -

~RAT

LLU

0^^^~~

~~Figure 4-3. Twin engine rate of climb (Sheet 5 of 24)

SEP O5 0Z2D.O 52OBB

Rev. Rev. 99

23 23


FAA APPROVED



70 KIAS HEATER ON


1 2 3

0

20000

1. CUMB SPEED IS 60 KIAS

2. RATE OF CLIMB WILL DECREASE 275 FT/MIN


18000

6

16000

14000

_=__--

4 0

.

12000

0000

2000 2400 200 3200 3600

4000 2000

0

400

800


Figure 4-3. Twin engine rate of climb (Sheet 6 of 24)

24

Rev. 9

$IEP0 5JfO

^


BHT-412-FMS-46.3


TAKEOFF POWER

70 KIAS


HEATER OFF


0



20000

i6ililiiii

ilhilHi.ii i'im

illl

lliliiilil

18000--------16000

--

i-

5

--

~

~L

~'

-'~..

-I

, 14000 2-

,,

_,\\ ~

\.

~1

\.

_

VN^^S^:

40

o

12000

0

I.8000------

---- RAE

6000

4000

',\

F CIM

- FT/NLJ

2UD -E 5----^ ~~~~~~SQ-

Rev.9

25

4000

0

400

800

Figure 4-3.



SEP 0 5 2000

Rev. 9

25

T N I R P N O R 0 I O T F C U T O OD F N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

FAA APPROVED


TAKEOFF POWER ENGINE RPM 100V% GENERATOR 150 AMPS (EA.)

70 KIAS HEATER ON


0

20000


1 2 3

T

u.Il,h

18000

14000

.

12000

ll

:lII l

l

ll U

l :

I

61i

---

16000 A "

w


Ns1-'"' r

5

-

40

-

_

-

8000 i- 10000

-

--

--

V000

-

-

r'

4000

0

400

800


Figure 4-3.

26

Rev.

9

0

SEP O 0 2 00



BHT-412-FMS-46.3


70 KIAS

TAKEOFF POWER


HEATER OFF



RATE OF CLIMB -

20000

,,, mllltlrlU ±lull ~~~~~~~~~~20000~~ ~ ~ '.ll'uhl

'" iinlil

M/SEC

7 8 9 1011 12 131415161718

3 4 5 6

1 2

0

II

llrJI 1

-

ii, Hal mLIulIl

6

18000-

16000 _< 14000

,.

R

\

\

-_ ,

__

-:

5

'

0~~~~o

I-

40

3

R/)

pI 0 0

4

01

12000

10000

*

--

>-¢-------

I

O

-

\M

I I I I e nginerte of cli (e 9 o 24)o On.A SEP00 400 800 1200 1600 2000 2400 2800 3200 3600

0^^r~~

~~Figure 4-3.

.0

_

RATE OF CLIMB -

FT/MIN


SEP 052000

Rev. 9

27


FAA APPROVED



70 KIAS HEATER ON


0



1 2

20000

i

iii

18000

-

16000

-

----

-n-

…-

§

9 28 12000

I-

U)

8000

1

--------------

Of,

0

400

800

Figure 4-3.

28

Rev. 9

SEP 0 5 Zfl

:-2 , i




FAA APPROVED


70 KIAS

TAKEOFF POWER

HEATER OFF




Fty~ RATE OF CLIMB - M/SEC 5 6 7 8 9 1011 12 131415161718

tb

0

1 2 3 4

- 6

20000_________

----------------

18000

16000

|

\l

.l-I

14000

3

II\

_

_

-

n

I II;I-I -I

-

40

12000

I-.

1? loooo

\

3 C1

…E'OF… IM-

,:,,,RA

U)

-

F/

----

8000

4000-----

2000-.-

000-,-----" 400

800

1200 Nil 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

I^^^-~

~

Figure 4-3.

0

FT/MIN


520 SEP aO

Rev. 9

29


FAA APPROVED

TWIN ENGINE RATE OF CUMB GROSS WEIGHT 11,900 LB (5398 kg)


70 KIAS HEATER ON


0


RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 9 1011 12 131415161718 u

20000,[],, .[,lI[Hlulilll[[[[htl.Hl[dElHlillIlllIIl

l-6

18000

16000o

5

,

14000o

L

,

\o

_ ^

12000__X

L

.10000

\

--

o

30i

6R000ev. 90 SP O 0 400 800

-- FT/MIN

-1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB - FT/MIN

Figure 4-3.

Rev.9

I

2III

RATEOL-

30

~,

…

0

30

0

______-__

t _

00

8000

1

40

N

~-

climb (Sheet 12 of 24) rate of climb Twin engine engine rate

SEP05 0DW

T N I R P N O R 0~ I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

FAA APPROVED


70 KIAS


HEATER OFF


0

1 2 3 4

20000

1

--

16000

I--I-

14000


RATE OF CLIMB - M/SEC 5 6 7 8 9 101112 131415161718

I

-

ILt

-

--

S

h

C

-------

-

--

)

---

40

.L

12 0 0 0

------

R

OI

5 t

"' 6000 11

V

R 9 -J 3

46000

4000

_o

I

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

Figure 4-3.

FT/MIN


SEP052000

Rev.9

31

T N I R P N O R I O T dl6 Q, F ,n C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

FAA APPROVED



70 KIAS HEATER ON




0 1 2 3 4 5 6

7 8 9 1011 12 131415161718 o 6

20000 18000

-V--

\

L.

-

" ,I

0

40

16000

14000

-

20_

_0

\

_

_ - Tl1 _ _

__Fiur

/~-

0-

3

------

--------

1i

11

///

-----------

6000

°

-----

ia:

w

\4-3. ,0v.9

uj 12000--

10000

3

-

-

-2-

3

"I

4000--

-

-

800

1200 1600 2000 2400 2800 3200 3600

-

/

j

--

-

-I

2000

0

400


Figure 4-3.

32


Rev. 9 SEP 0o5Z0Oa

0

T N I R P N O R 0 I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-46.3

FAA APPROVED


70 KIAS


HEATER OFF



0

1 2

20000

18000


RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 101112 131415161718 i i m0

-

"

16000

--

_

14000

40

0

U'. . 12000 UJ4

10000 :"

--

-----

000-----_000

0

400

800

Figure 4-3.



SEP 052Q

Rev.9

33


FAA APPROVED



70 KIAS HEATER ON


1. CUMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FTIMIN

RATE OF CLIMB - M/SEC 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

20000

18000

-

5

16000

34

,

12000

SEP

Rev49

Q5M0

-

5 10000

3

--

Uj~~

~

~

~~~~~~~I

l

. 8000-

2000 4000

400

800

Figure 4-3.

,ev. e349 Rev.

n

/I fl 1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB

34

/

--

2000

01 0

-

vwIl

-

FT/MIN


SEP 05200

0


BHT-412-FMS-46.3



70 KIAS


HEATER OFF


0

1 2



20000

,6hlIiil n

~6o~loo _ 16000

_

6

l nln,'nlI i.InWbni

-

.,, \,n'

V §

514000

0

812000

<

*

i. 6

A

-000

C

-

v--

a uT

2000

4000

0

400

800

Figure 4-3.



SEP 0520

Rev. 9

35

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L N ooo O FAA APPROVED

BHT-412-FMS-46.3


70 KIAS


HEATER ON




RATE OF CLIMB -

0

20000

3 4 5

1 2

6 7

1111

i'l!

18000

-

:

-

RAT

8000 28000

o.

6000

3-

240

O

20

CIM

3

3, :

0

-

-

12000

n,

611

lilI'lIU:il'l

.

16000

,

M/SEC

8 9 10 11 12 13 14 15 16 17 18

- 2

^

4000 2000

-

0

400

800

1200 1600 2000 2400 2800 3200 3600 RATE OF CLIMB -

Fr/MIN


36

Rev. 9

SEP 052000


BHT-412-FMS-46.3



70 KIAS


HEATER OFF


0



20000 lullll

ul

18000

1\-

16000

'"~--°…_

l

,,ulll,,llllI

lhl1l

_"-.

V-

Id

I 14000\

M

_…_ \^_

__…________…__-40

,J 12000

-L o-.

- 10000

-

--

--

-3

0

8000

-i

0Q

000

-

-

-

V)

l---

4|-

4000

2000

0

400

800



SEP 052Os

Rev.9

37


FAA APPROVED



70 KIAS HEATER ON


0


RATE OF CLIMB - M/SEC 1 2 3 4 5 6 7 8 9 1011 12 131415161718

6m

20000

18000-

p10000_ _ _ _

16000

\g

_

_

_ _ __X ----

__ -3°5

w

a

I

14000

4

-----

10000

----------

3

<

8000

0

400

800

Figure 4-3.

38

Rev. 9

0



EP 5 nnn


BHT-412-FMS-46.3



70 KIAS


HEATER OFF


0



6

20000

4------2 -000 __ _

1 18000

16000

5

I

-5

^

I-

^ _ _

14000

_

_ __

_

_

_

_

_

_

_

_

I-£

_

u, 12000

1.10000

40

81030

rtA

8ooo

-6ooo-

)w 60 2000

----4000------

2000

<

---

0

400

800



SEP 052000

Rev. 9

39


FAA APPROVED

TWIN ENGINE RATE OF CUMB GROSS WEIGHT 11,000 LB (4990 kg)


70 KIAS HEATER ON


0

1 2

20000


RATE OF CLIMB - M/SEC 3 4 5 6 7 8 9 1011 121314 15161718 1L l i li l il l J 11 L

18000__

_

16000I

__

6000\n\X….__

L.

-----

-

14000

-

O

0 o

12000 o000

--

8000 2

t

0

400

800

Figure 4-3.

Rev.9

4

lil

-

1200 1600 2000 2400 2800 3200 300 …-

RATE OF CLIMB -

40

JX

__a___

FT/MIN


SEP 052 000


BHT-412-FMS-46.3



70 KIAS


HEATER OFF


I^^w~

0


~~RATE OF CLIMB - M/SEC

1 2

3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18

20000 "" '' ' i

'l

IllilJI

mll

l

l

ll nl

l

l

18000

16000

,

--------

-

14000

40

12000

i10000

----

-

o

o

4-30

…

i.i

I-

8000 6000

.

20

<

30

-2

---

4000



SEP 052000n

Re0.

0


FAA APPROVED



70 KIAS

ENGINE RPM 100%

HEATER ON




RATE OF CLIMB -

20000

2 14000

'\

0

4

ili

ii

M/SEC

5 6 7 8 9 10 11 12 13 14 15 16 17 18

O 1 2 3 4

llll ilill

'il l ili ulil

iliuii

ulIllI

l'

il li il

uil

-

4_

_

_

_0_

_

_

-4 0

S-5-0

12000 9

3

000

8000----…----------

4000--

--

2000

0

400

800

0


Figure 4-3.

42

Rev. 9


SEP 052000


BHT-412-FMS-46.3





~~RATE OF CLIMB - M/SEC

D^^-~

-4 -3 -2 -1 20000 L'. 18000

16000

,.

0

1 2

4

5

6

Fgr4---

Sin

egin

8

9 10 11 12 lj .LI 6

o

--

6000

5,

--

14000

-

--

-

-

-

-

-

30I.-

---

--------

\\

*~\ \ \ -- ----

\\\-\'~\

2000o0 -800

7

mj

4ooo - -- -- - --\ 4 6000 ----------

*

3

-z

wI

o-


---

-

-400

0

400

800

RATE OF CLIMB

Figure 4-4.

1200

-

--

1600

2000

0^

2400

FT/MIN

Single engine rate of climb chart (Sheet 1 of 6)

SP

05 2000

Rev. 9

43

SEP 052000

Rev. 9

43

T N I R P N O 0 R I O T F C U T O OD · N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-46.3

SINGLE ENGINE RATE OF CLMB GROSS WEIGHT 8000 LB (3629 kg)





RATE OF CLIMB - M/SEC -6 -5-4-3 -2-1 0 1 2 3 4 5 6

20000

_L<1i

18000

-'-

7

8

9 10

i

-

16000

:

-

LaJ

2W

1400

--

-

_

1

,, 12000 g

0^

2000

-6

200 -800

Rev. 99 Rev.

"

"

--

I

04 5

-^

-

,

-

Figure 4-4.

44 44

-

RT,--.-

-j 10000

In

-"

^se~~s \\Y

--" -_

-00

00

800

200

600


SEP 52000 SE 0 05 2000

2000

4


FAA APPROVED

SINGLE ENGINE RATE OF CUMB GROSS WEIGHT 9000 LB (4082 kg)

70 KIAS






M/SEC

RATE OF CLIMB -

-8 -7 -6 -5 -4 -3 -2 -1

20000

iL

E il

IE

11

0

1 2

4

3

i lll

I

6

5

il

II

8

7

ii1i6

TWIN EGI E

16000 16000 ABSOLL CEIUN-------INCTE 14000

,j

5

_

-

12000

RTOFCM

\

-0T4

03 00nasa

810000

œ

8000-

\ \

n \

Li

-

6000 8

-'0

V)

---

LU I-

0

4000

2000

-1600

---

-1200

--

-800

--

-400

--

0

RATE OF CLIMB -

Figure 4-4.

400

--

^

^

-

800

1200

1600

FT/MIN


SEP 05 2000

Rev. 9 Rev. 9

45

45

T N I R P N O R 0 I O T F C U T _6x O OD N R E P E E R R o_a= T I I _ _ _SC \ I o N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-46.3

SINGLE ENGINE RATE OF CUMB







-10-9 -8-7 -6-5-4-3-2 LIII ,l ,ll ,l, I 20000 18000-------

0

1

,I,l

l =

2

3

4

5

6

6

IUIU

NGIIE_CF

0TWIN

-

16000

,

-1

-----

_

-

-_)

w 12000

E i

LLL

8000-10-

40

--------

L

-4---0

4 12000------

-2000-1600-1200

-800

RATE

-400

OF CLIMB -

0

400

800

FT/MIN

Figure 4-4. Single engine rate of climb chart (Sheet 4 of 6)

46

Rev. 9

SEP 052000

0


FAA APPROVED




1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FTIMIN


RATE OF CLIMB - M/SEC -10-9 -8-7-6-5-4 -3-2-1 0 1 2 _ llttl i l cimii I 20000 I lr I II

--TWIN

18000

3

4

5

-6

ENGIME - I

5

16000 AB OLJTE CEILIN(

14000 14000 --

U

6

w

-

4

120000

..

,

'"-

_

8000

\

-

-

3

Uj

-

w

6000

4000

----

---

-

2000

~*

0

---

0--jib---

-2000 -1600 -1200

-800

RATE

Figure 4-4.

-400

OF CLIMB -

0

400

800

1200

FT/MIN


SEP 05 2000

Rev. 9

47

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F Y V L N 0 \-6_ O _ -0 BHT-412-FMS-46.3

FAA APPROVED






RATE OF CLIMB - M/SEC -10-9 -8-7-6-5-4 -3 -2 -1 0 1 2

20000 Il,,,l,,I

,,,l,,,,.1

ii

' a LLI iii

3

4

Il

5

6

,,lI

----

18000

14000 mV' " ABELLATE CEIINGC

,/I

4 12000 -

4

--

0

l-

w

10000

En

- 3 0-j

--

Iw\,\

i"

a

- -

-

\ \ \0

6000

!

Li

-

Li

4000-

------

-2000 -1600 -1200

-800

-400

RATE OF CLIMB -

0

400

--800

Fr/MIN

Figure 4-4. Single engine rate of climb chart (Sheet 6 of 6)

48

Rev. 9

SEP 05 ZOO0

1200

0


MODEL


FOLDINGSTEP 412-899-287

CERTIFIED 1993 25 OCTOBER

This supplement shall be attachedto the Model 412 Flight folding step has beeninstalled. Manualwhen the 412-899-287 Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and PerformanceData not containedIn this supplement,consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1993 INC. BELL* HELICOPTER TEXTRON. ANDBELLHELICOPTER LTD. CANADA OFTEXTRON ADIVISION RESERVED ALLRIGHTS

POST OFFICE BOX 482

Inc. of Textron A Subsidiary


25 OCTOBER 1993


NOTICE PAGE

Additional copies of this publicationmaybe obtainedby contacting: CommercialPublicationDistributionCenter Bell HelicopterTextron P. O.Box 482 Fort Worth,Texas76101-0482

NP


LOG OF REVISIONS


0 ..............

25 OCT 93

LOG OF PAGES

REVISION NO.

PAGE

FLIGHTMANUAL

Title - NP .............................. A - B ..................................

REVISION NO.

PAGE

1-2...................................0

0 0

MANUFACTURER'S DATA

NOTE



FAA APPROVED


Original

...........

0 ..............

25 OCT 93

APPROVED:

MANAGER


B

T N I Section R P N O R I O T F C U T O OD N R E P E E R Section2 SCR N G O N I R W O E I F V NLY O BHT-412-FMS-47

If necessary, to return empty weight CG within allowable limits.


The 412-706-004 Emergency float kit shall not be installed in conjunction with the folding step.

WEIGHT -

CG LIMITATIONS

Actual weight change shall be determined after kit is installed and ballast readjusted,

NOTE

6.

Folding step -

EXTERIORCHECK

2.

FUSELAGE

SIDE

-

CABIN

Folding step - Stowed (up).

FUSELAGE -

SIDE

After passenger loading/ unloading, stow step (up).

LEFT

CABIN RIGHT

Stowed (up).

T N Section 3 I R P N O R I O T F C Section 4 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-47



BHT-412-FMS-48.1 & 48.2

T N I R P ROTORCRAFT N O R FLIGHT MANUAL I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MODEL

SUPPLEMENT

ENGINE NO. 2 GOVERNOR TRIM SWITCH TB 412-93-118 33001 36001 -

33213 36019

CERTIFIED 28 JULY 1994

This supplement shall be attached to the Bell Helicopter Model 412 Flight Manual (BHT-412FM-1 and BHT-412-FM-2) when engine #2 governor trim switch has been installed per TB 412-93-118. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1994 BELL HELICOPTERINC

ANDBELLHELICOPTER TEXTRON. A DIVISIONOF TEXTRONCANADA LTD.

ALLRIGHTSRESERVED

A Subsidiary of Textron Inc. POST OFFICE BOX 482 · FORT WORTH. TEXAS 76101

28 JULY

1994

BHT-412-FMS-48.1& 48.2

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O NOTICEPAGE


Manufacturer's Data portion of this supplement is proprietary to Beli Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron, Inc.


FAA APPROVED

BHT-412-FMS-48.1& 48.2

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOGOF REVISIONS


0..............28

JUL 94

~LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP .............................. A - B .................................. 1 - 2 ...................................

REVISION NO.

PAGE

MANUFACTURER'SDATA

0 0 0

3/4 ......................................

0

NOTE


BHT-412-FMS-48.1 & 48.2

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O LOGOF APPROVED REVISIONS


0..............

28 JUL 94

APPROVED:

MANAGER


FAA APPROVED

BHT-412-FMS-48.1 AND 48.2

T Section 1 RIN P N O R I O T F C U T Section 2 O OD N R E P E E R R C S N G O N I R W O E I F V NLY O No change from basic manual.

BEFORE TAKEOFF Throttles frictions.

-

TAKEOFF

Full open.

Adjust

Area -

Clear.

NOTE

RPM INCR/DECR switch - Minimum beep (DECR for 4-5 seconds).

RPM INCR/DECRswitch trim (-2 for 4-5 seconds). NR-

Minimum

Check 95% or greater.

RPM INCR/DECRswitch - Adjust to obtain matching torque at 100% NR. Flight instruments operation and set.

-

Check

As collective is increased, it may be necessary to rematch engine torque prior to reaching hover.

RPM INCR/DECRswitch - Adjust to obtain matching torque or ITT, as required, and 100% NR. Hover power - Check torque required to hover at four feet skid height.

BHT-412-FMS-48.1AND 48.2

FAA APPROVED

T N I Section 3 R P N O R I O T F C U T Section 4 O OD N R E P E E R R C S Section 5 N G O N I R W O FSection6VIE LY N O No change from basic manual.




BHT-412-FMS-48.1AND 48.2

T Section1 RIN P N O R I O T F C U T Section 2 O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA


MANUFACTURER'S DATA

_

The +2/-2 switch allows the pilot to match engine performance and improve total engine power available.

ON

INCREASEENGINE RPM(BOTHPOWERSECTIONS)

+2-

2

ARMED F

G RET IDLESTOP

OFF S START

ENG

INCR

DECREASE POWER SECTIONNO. 2 RPM

EXT

INCREASEPOWER SECTIONNO. 2 RPM

UP

ENG

ENG

SRCHLT

DECR

DECREASEENGINE RPM IBOTHPOWERSECTIONS

412-FMS-48.1 48.2-2-1

Figure 2-1. Engine RPM INCR/DECRswitch. 3/4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O This supplement shall be attached to Model 412EP Flight Manual when Honeywell Primus 700 Weather Radar kit has been installed.


TEXTRON. ANDBELLHELICOPTER A DIVISIONOF TEXTRONCANADA LTD.

RESERVED ALLRIGHTS

05 OCTOBER 1994


NOTICEPAGE


is Data portion of this supplement Manufacturer's proprietary to Bell Helicopter Textron, Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron, Inc.

Additional copies of this publication may be obtained by contacting: Commercial Publication Distribution Center Bell Helicopter Textron P. O. Box 482 Fort Worth, Texas 76101-0482

NP


LOGOF REVISIONS


0..........

05 OCT 1994

LOG OF PAGES

MANUFACTURER'S DATA

FLIGHT MANUAL

PAGE

Title - NP .............................. A - B .................................. i/ii ....................................... 1 - 2 ...................................

REVISION NO.

PAGE

REVISION NO.

3/4......................................0

0 0 0 0

NOTE



BHT-412-FMS-49.4



0 ...........

05 OCT 1994

MANAGER


B

BHT-412-FMS-49.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O GENERAL INFORMATION

The Honeywell Primus Weather Radar kit (412-705-034) consists of a digital indicator, receiver/transmitter, antenna, wiring and hardware necessary for installation. The digital indicator is mounted in the instrument panel and contains all controls necessary to operate the radar. The antenna and receiver/ transmitter are located in the nose of the helicopter. The weather radar is used for weather detecting and analysis and for ground

mapping. The weather mode detects and locates various types of storms along the flight path of the helicopter, and gives the pilot a visual indication of storm intensity. After proper evaluation, the pilot can chart his course around these storm areas. The mapping mode enables the pilot to identify coastlines, hilly or mountainous regions, and cities. The weather radar performs only the functions of weather detection and ground mapping and shall not be used, nor relied upon, for proximity warning or collision avoidance.

FAA APPROVED

T N I Section1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-49.4

1-3. TYPES OF OPERATION 1-3-C. GROUNDOPERATION

Radar system shall not be operated on ground or in a hover when personnel are in the danger area shown in figure 1-1. Radius (R) of danger area is relative to size of antenna. Radar system shall not be operated within 100 feet (30.5 meters) of any fueling operation.

Radar system shall not be operated on ground anytime a large metallic object is forward of helicopter nose, within 65 degrees of centerline and at a distance of less than 100 feet (30.5 meters). WARNING

Honeywell Primus 700 Weather Radar system performs only functions of weather detection and ground mapping. It shall not be used for ground proximity warning or collision avoidance.

ANTENNA

RADIUS

412-FS-49.4.1.1

Figure 1-1. Danger area

1

T N I R P N O R I Section2 O T F C U T O OD N R E P E E R R C S Section 3 N G O N I R W O E I F Y V L N Section 4 O FAA APPROVED

BHT-412-FMS-49.4

1-6. WEIGHT AND CENTEROF GRAVITY

If necessary, to return empty weight CG to within allowable limits.

Actual weight change shall be determined after kit is installed and ballast readjusted,

Refer to Honeywell Publication No. 281146-66-00Pilot's Manual.

2-7. BEFORETAKEOFF Weather radar - Standby (STBY)

2-9. IN-FLIGHTOPERATIONS Weather radar -

On


Weather radar - Standby (STBY)



T N I Section5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

Actual weight change shall be determined after kit is Installed and ballast readjusted,

BHT-412-FMS-49.4

if necessary, to return empty weight CG to within allowable limits.

3/4

T N I R P N O ROTORCRAFT R I O T FLIGHT MANUAL F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-53.3& 53.4

MODEL

SUPPLEMENT

PT6T-3DE ENGINE (30 MINUTE OEI RATING) 412-706-050 32020 -

36086

AND 36087 & SUB

CERTIFIED 29 MARCH 1996

This supplement shall be attached to Model 412/412EP Flight Manual,when PT6T-3DEENGINE(30 MINUTEOEI RATING)kit has beeninstalled. Information contained herein supplements information fo basic Flight Manual.For Limitations, Procedures,and Performance Data not contained in this supplement, consult basic Flight Manual.

COPYRIGHT NOTICE BELL0 HELICOPTER INC. AND BELLHELICOPTER TEXTRONINC. A DIVISION OF TEXTRON CANADALTD. ALL RIGHTSRESERVED

REISSUE

21 JANUARY

1999


NOTICEPAGE


NP

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-53.3& 53.4

LOG OF REVISIONS

Original ...................... 0 .............. Revision ..................... ...............

29 MAR 96 23 APR 96

Reissue......................

................

21 JAN 99

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

REVISION PAGE NO A - B ................................................... 1 i/ i .............................................................

Title .......................................................... 1 NP........................................................... 0

0

1 - 1 4 ...................................................... 0 15/16 ........................................................ 0

NOTE



BHT-412-FMS-53.3& 53.4


Original ....................... Revision......................

................ 29 M AR 96 ................ 23 APR 96

APPROVED:

ANAGER


Reissue ...................... 0 ...............

DATE:

21 JAN 99

JAN21 1999


GENERAL INFORMATION

The PT6T-3Dengine has been redesignated as PT6T-3DEto offer a 30 Minute OEI rating. This 30 Minute OEI rating represents improved OEI capabilities, but maximum continuous OEI performance is reduced from that presented in the Basic Flight Manual.


FAA APPROVED

1-1. INTRODUCTION

PT6T-3DEengine has been revised to offer a 30 Minute OEI rating.

Maximum Continuous and Takeoff OEI Operation 30 Minute OEI

103.2% 103.2 to 109.2% 105.7% 109.2%

1-11. AMBIENT TEMPERATURES

21/2 Minute OEI

The maximum sea level ambient air temperature for operation is 48.9°C (+120°F).


1-13. POWER PLANT

Pratt and Whitney Aircraft of Canada, Ltd. PT6T-3DE. NOTE

Operation in an OEI range is intended for emergency use only when one engine becomes inoperative due to an actual malfunction.

For Power Assurance check only, engine torque value of 59 to 70% within OEI range may be utilized.

Anytime an engine is operated in an OEI range, an entry shall be made in the helicopter logbook detailing the extent of operation in excess of twin engine takeoff power limits. This does not apply to approved ITT limits for starting or Power Assurance Check.

1-13-A. GAS PRODUCERRPM Continuous Operation

61 to 103.2%

Continuous Operation Maximum Continuous Maximum Transient (5 seconds above 810°C) OEI Operation 30 Minute OEI

300 to 810° C 810°C

21/2 Minute OEI

940°C 810 to 940°C 855°C 940°C

Maximum for Starting (2 seconds above 960°C)

1090°C

1-13-D. ENGINETORQUE

Continuous Operation Maximum Continuous OEI Operation 30 Minute OEI

5 to 59% 59% 59 to 81%

2'/2 Minute OEI

81%

77%

1-20. INSTRUMENT MARKINGS AND PLACARDS

Refer to Figure 1-1 for instrument range markings and Figure 1-2 for placards and decals. 1

BHT-412-FMS-63.3 AND63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Figure1-I. Instrument markings


FAA APPROVED

TWIN & MAX CONT OEI 103.2% 2.5 MINOEI 109.2% 30MINOEI 105.7%

412FS53.4-1-2

Figure 1-2. Placards and decals

3

T Section2 N I R P N O R I O T F C 3 Section U T O OD N R E P E E R Section 4 SCR N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-53.3 AND 53.4



4-1. INTRODUCTION

Performance data presented herein are derived from engine manufacturer's specification power for PT6T-3DE engine less installation losses.

4-2. POWER ASSURANCE CHECK

Power Assurance Check charts (figure 4-1) are provided to determine if the engines can produce installed specification power. A power assurance check should be performed daily. Additional checks should be made if unusual operating conditions or indications arise. The hover check is 4

performed prior to takeoff, and in-flight check is provided for periodic in-flight monitoring of engine performance. Either power assurance check method may be selected at the discretion of the pilot. It is the pilots responsibility to accomplish the procedure safely, consider passenger load, terrain being overflown, and the qualifications of persons on board to assist in watching for other air traffic and to record power check data.

If either engine does not meet the requirements of the hover or or the inflight power assurance check, published performance may not be achievable. The cause of engine power loss, or excessive ITT should be determined as soon as practical. Refer to Engine Maintenance Manual.


BHT-412-FMS-53.3 AND 53.4


Refer to Figure 4-2 for Single engine rate of climb - 30 minute power. Refer to Figure 4-3 for Single engine rate of climb maximum continuous power.

5


BHT-412-FMS-53.3 AND 53.4

MODEL412 POWER ASSURANCE CHECK- HOVER PT6T-3DEENGINE

HEATER/ECU- OFF.


THROTTLES: TEST ENGINE- FULL OPEN, FRICTIONED. OTHERENGINE- FLIGHT IDLE. NII RPM -


7%.

COLLECTIVE PITCH - INCREASE UNTIL LIGHT ON ° SKIDS OR HOVERING. DO NOT EXCEED 810 ITT, 103.2% NI RPM, OR 70% TORQUE.

IF EITHER ENGINE EXCEEDS ALLOWABLE ITT AFTER STABILIZING FOUR MINUTES, PUBLISHEDPERFORMANCE MAYNOT BEACHIEVABLE.CAUSESHOULDBE DETERMINED ASSOON AS PRACTICAL.

STABILIZEPOWERONE MINUTE. THEN RECORDPRESSURE ALTITUDE,OAT,ENGINETORQUE AND ITT. ENTERCHARTAT INDICATEDENGINETORQUE.MOVE UP TO INTERSECTPRESSUREALTITUDE. PROCEEDTO THE RIGHT THEN MOVE UP TO INTERSECTOUTSIDEAIR TEMPERATURE, TOREAD VALVEFORMAXIMUM ALLOWABLEITT.

550

MAXIMUM ALLOWABLEITT - °C 600 660 700


412FS53.4-4-1-1

Figure 4-1.

Power assurance check - (Sheet 1 of 2)


BHT-412-FMS-53.3 AND 53.4

MODEL 412 POWERASSURANCE CHECK - IN-FLIGHT PT6T-3DE ENGINE

ESTABLISHLEVELFLIGHTABOVE 1000 FEETAGL AIRSPEED- 100 KIAS (OR VNE.IF LESSI.

ENTERCHARTAT INDICATEDENGINETORQUE,MOVE UP TO INTERSECTPRESSURE ALTITUDE. PROCEEDTO THE RIGHT TO INTERSECT OUTSIDEAIR TEMPERATURE, THENMOVE UP TO READVALUEFORMAXIMUM ALLOWABLEITT.

HEATER/ECU- OFF.


THROTTLES: TESTENGINE- FULL OPEN.FRICTIONED

OTHER ENGINE - DECREASE SLOWLY UNTIL TEST ENGINE TORQUE IS WITHIN TEST RANGE. DO NOT EXCEED 810 C ITT, 103.2% NI RPM, OR 70% TORQUE.

NII RPM - 97%.

CHECK USINGOTHERENGINE. REPEAT

IF EITHER ENGINE EXCEEDS ALLOWABLE ITT AFTER STABILIZINGFOUR MINUTES, PUBLISHEDPERFORMANCE MAY NOT BEACHIEVABLE.CAUSESHOULDBEDETERMINED AS SOONAS PRACTICAL.

STABILIZEPOWER ONE MINUTE IN LEVEL FLIGHT,THEN RECORDPRESSUREALTITUDE,OAT. ENGINETORQUE AND ITT

500

550

MAXIMUM ALLOWABLEITT - °C 750 700 660 600

800


412FS53.4-4-1-2

Figure 4-1.

Power assurance check - (Sheet 2 of 2)

7


BHT-412-FMS-53.3 AND 53.4



70 KIAS

GROSS WEIGHT7000LB (3175KG)


WITHALL DOORSOPEN OR REMOVED: 1. CLIMBSPEED IS 60 KIAS 2. RATE OF CLIMBWILL DECREASE275 FT/MIN

RATE OF CLIMB - METERS / SECOND

-4

-2

0

2

4

6

8

10

12

-

6

18000 16000

412FS53.4-4-2-1

Figure 4-2. Single engine rate of climb - 30 minute power (Sheet 1 of 6)


BHT-412-FMS-53.3 AND 53.4


30 MINUTE POWER GROSS WEIGHT8000LB(3629 KG) ENGINE RPM 97% HEATER OFF GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED WITHALLDOORS OPENORREMOVED: 1. CLIMBSPEEDIS60KIAS

2. RATE OF CLIMBWILL DECREASE275 FT/MIN


-6

-4

-2

0

2

4

6

412FS53.4-4-2-2

Figure 4-2.

I

Single engine rate of climb - 30 minute power (Sheet 2 of 6)

9


FAA APPROVED

SINGLE ENGINE RATE OF CLIMB GROSSWEIGHT9000 LB (4082 KG)

POWER

30 MINUTE

ENGINE RPM 97% GENERATOR150 AMPS


WITHALL DOORSOPEN OR REMOVED: 1. CLIMB SPEED IS 60 KIAS 2. RATEOF CLIMB WILL DECREASE275 FT/MIN


-8

-6

-4

-2

0

2

0

400

4

6

8

18000 16000

-1600 -1200 -800 -400

800

1200

1600

RATEOF CLIMB- FEET/MINUTE

412FS53.4-4-2-3



FAA APPROVED

SINGLE ENGINERATEOF CLIMB

70 KIAS GROSSWEIGHT10,000LB(4536KG) 30 MINUTE POWER HEATER OFF ENGINE RPM 97% INOPERATIVEENGINE SECURED GENERATOR 150 AMPS 1. CLIMBSPEEDIS 60KIAS OPENORREMOVED: WITHALLDOORS 275FT/MIN 2. RATEOFCLIMBWILLDECREASE RATE OF CLIMB - METERS/ SECOND

-8

-6

-1600 -1200

-4

-2

0

2

-800

-400

0

400

4

800

6

1200

8

1600


412FS53.4-4-2-4

Figure 4-2.


11


FAA APPROVED


70 KIAS GROSSWEIGHT11,000LB(4990KG) 30 MINUTE POWER ENGINE RPM97% HEATER OFF INOPERATIVE ENGINE SECURED GENERATOR 150 AMPS WITHALL DOORSOPENOR REMOVED: 1.CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMBWILL DECREASE275 FT/MIN


-8

-6

-4

-2

0

2

4

6

8

6

18000

5

16000


412FS53.4-4-2-5

Figure 4-2.


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 18000 16000

Figure 4-2.

Single engine

I

rate of climb - 30 minute power (Sheet 6 of 6)

13


FAA APPROVED


MAXIMUMCONTINUOUSPOWER ENG- 97% RPM (N2) GENERATOR 150 AMPS


WITHALL DOORS OPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMB WILL DECREASE 275 FT /MIN

20,000

18,000

16,000

I

Figure 4-3.

Single engine rate of climb - maximum continuous

power (Sheet 1 of 6)


BHT-412-FMS-53.3 AND 53.4


GROSS WEIGHT 8000 LB (3629 kg) MAXIMUM CONTINUOUSPOWER 70 KIAS ENG - 97% RPM (N2) HEATER OFF GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED WITHALL DOORSOPENOR REMOVED 1. CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMB WILL DECREASE 275 FT /MIN

Figure 4-3.



15

BHT-412-FMS-53.3

AND 53.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPRO


MAXIMUM CONTINUOUSPOWER ENG- 97% RPM (N2) GENERATOR 150 AMPS


WITH ALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT /MIN

412FS53.4-4-3-3

I

Figure 4-3.



T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 20.000

18,000 -TWIN ENGINE MC P ABSOLUTE CEILING

Figure 4-3.



17


FAA APPROVED

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT11.000 LB (4990 kg)

MAXIMUM CONTINUOUSPOWER ENG- 97% RPM(N2) GENERATOR 150 AMPS


WITHALL DOORSOPENOR REMOVED I CLIMB SPEEDIS 60 KIAS 2 RATEOF CLIMB WILL DECREASE 275 FT /MIN

-6000

- 5500

-5000

-4500

RATE OF CLIMB - (METERS/SECOND)

I

Figure 4-3.






19/20


ROTORCRAFT FLIGHT MANUAL

SUPPLEMENT EMERGENCY FLOATS WITH AUTOMATIC RETRACTING PASSENGER STEP 412-706-004 S/N 36122, 36123 S/N 36125 AND SUB CERTIFIED 15 APRIL 1996

This supplement shall be attached to Model 412/412 EP Flight Manual when 412-706-004Emergency Floats have been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter i

COPYRIGHT NOTICE COPYRIGHT 1997 BELL* HELICOPTER TEXTRON INC. ANDBELLHELICOPTER TEXTRON INC ADIVISION OFTEXTRON CANADA LTD ALLRIGHTSRESERVED

Subsdiary of TextronInc POST OFFICE BOX 482 * FORT WORTH, TEXAS 76101

REISSUE -

31 MAY 1997


NOTICEPAGE



LOG OF REVISIONS

Original ........... Reissue...........0

0 .............. .............

15 APR 96 31 MAY 97

LOG OF PAGES

PAGE

REVISION NO.

Title - NP .............................. AB ..................................

0 0

i/ii .......................................

0

PAGE

1 - 4 ................................... 5/6 ......................................

REVISION NO. 0 0

NOTE



LOG OF FAA APPROVEDREVISIONS


0 ............. 0.............

15 APR 96 31 MAY 97

APPROVED:

MANAGER

ROTORCRAFTCERTIFICATIONOFFICE FEDERAL AVtATION ADMINISTRATION FT. WORTH, TX 76193-0170

DATE:


GENERAL INFORMATION

Emergency floats will allow helicopter to land In water during an emergency situation. Kit consists of six skid mounted float bags, four water activated float switches that are mounted on underside of fuselage, a pneumatic bottle located in nose compartment, a collective mounted FLOATS switch, a pedestal mounted EMER INFLATION handle, and necessary hardware to complete installation. Kit also includes provisions for installing life rafts and vests.

Floats automatically inflate when FLOATS switch is in ARMED position and water activated float switches are submerged. If system or one float bag ruptures after inflation, check valves in each float bag will maintain float integrity. An instrument panel mounted EMER FLOATS caution light illuminates whenever FLOATS switch is in ARMED position. Placing FLOATS switch to ARMED position automatically retracts passenger steps to full stow. System security is protected by an EMERG FLOATS circuit breaker located on overhead console. EMER INFLATIONhandle provides for manual activation of emergency flotation system in event floats fail to inflate automatically.


FAA APPROVED


readjusted, if necessary, to return empty weight CG to within allowable limits.

for Emergency floats are installed assistance during emergency ditching and use is approved for VFR and IFR operation.

Installation of emergency floats does not change CG limits of helicopter.

Inflation of emergency floats using EMER INFLATIONhandle In flight Is prohibited Flight after landing with floats inflated is prohibited.

FLOATS switch shall be OFF and FLOATS caution light extinguished when operating over land. Emergency float safety pin shall be removed prior to flight over water.

1-5. CONFIGURATION

1-5-A. FLIGHTWITHOPTIONAL EQUIPMENTINSTALLED

Fixed passenger step shall not be installed concurrent with emergency floats. Retractable passenger step shall be stowed prior to flight over water.

When internal hoist is installed, hoist cable anti-chafing guard shall be installed on same side of helicopter as hoist.


Actual weight change shall be determined after floats are installed and ballast

1-7. AIRSPEED

WARNING

SEVERE NOSE UP PITCHING WILL OCCUR IF EMERGENCY FLOATS ARE INFLATED IN FORWARDFLIGHT OR DESCENT. REFER TO INADVERTENTFLOAT INFLATION IN FLIGHT (SECTION 3).

Maximum autorotation stowed):

airspeed (floats

105 knots below 10,000 feet Hp. 80 knots above 10,000 feet Hp.

Maximum forward speed for ditching: 33 knots in calm water. 15 knots in rough water.

Rate of descent should be reduced as low as possible upon water contact.

1-8. ALTITUDE

Maximum Hp for inflation of emergency floats is 10,000 feet. Helicopter operation above 10,000 feet is permitted provided FLOATS switch is in OFF position and FLOATScaution light is extinguished.

1

T N I R P N O R I O T F Section 2 C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-55.4

1-20. INSTRUMENT MARKINGSAND PLACARDS

FLOAT INFLATIONOR OPERATIONIN FORWARDFLIGHTIS PROHIBITED

1-22. FLOATARMING Floats shall not be armed: Over land. Above 60 KIAS. Above 600 feet AGL. Above 10,000 feet Hp.

(Located above pilot airspeed indicator.)

2-3. PREFLIGHTCHECK

STEP circuit breaker -

In.

EMER FLOATScircuit breaker - In.

2-3-B. EXTERIORCHECK

Verify FLOATS caution extinguished.

NOTE

STEP switch -

Ensure emergency floats have and inflation had periodic inspection.

Emergency floats -

Stowed.


EMER INFLATION handle safetied.

Observe both passenger steps extend, and STEP EXTND caution light illuminated. ARM.

Observe FLOATS caution light illuminated. Verify both passenger steps are fully retracted. Verify STEP EXTND caution light extinguished.

FLOATS switch -

(FOR FLIGHTOVER WATER)

Down and

Nitrogen bottle - Secured and pressure within allowable limits for Hp and ambient temperature as shown on chart decal, located on nitrogen bottle. Emergency floats safety pin - Remove for over water flight.

is

RAISE.

FLOAT switch -

Emergency float covers and supports Clean and secured.

light

STEP switch -

OFF.

STOW.

Verify passenger steps are fully retracted. Verify FLOAT and STEP EXTND caution lights are extinguished.

T N I R P N . O R I O T F C U T O OD N R E P E E R R C S N G O N I R 3 Section W O E I F Y V L . N O BHT-412-FMS-55.4

FAA APPROVED

BEFORE TAKEOFF

2-7.

2-10.

DESCENT AND

LANDING

2-7-A. BEFORE TAKEOFF OVER WATER

I STEP switch -

STOW.

2-10-A.

BEFORE LANDING OVER

STEP switch -

STOW.

FLOATS switch - ARMED; check FLOATS caution light illuminated.

FLOATS switch - ARMED when below 600 feet above water and airspeed of 60 KIAS

Check STEP EXTND light extinguished.

illuminated.

or below; check FLOATS caution light Check STEP EXTND light extinguished.

2-9.


STEP switch -

STOW.

2-10-B.

AFTER LANDING

FLOATS switch

-

OFF; check

FLOATS

caution light extinguished.

FLOATS switch - OFF upon reaching 600 feet above water and airspeed of 60 KIAS

or above, or when over land; check FLOATS caution light extinguished. NOTE

2-13.

BEFORE LEAVING

HELICOPTER

Emergency floats safety pin -

installed.

STEP EXTND warning light will remain illuminated with step in stowed position when both step and float switch are selected to OFF position.

3-7.

ELECTRICAL SYSTEM

NOTE

If EMER FLOATS circuit

3-7-A. FLOATS CAUTION LIGHT ON WITH FLOATS SWITCH OFF

If FLOATS caution light illuminates with FLOATS switch OFF, pull EMER FLOATS

breaker is

pulled, floats will not deploy automatically.

illuminated after pulling circuit breaker, land as soon as practical

circuit breaker.

3


3-7-B. STEPEXTNDCAUTION LIGHTON WITHFLOATSSWITCHIN ARMED If STEP EXTND caution light illuminates with FLOATS switch in ARMED, place step switch to STOW. If caution light remains illuminated, visually verify step(s) position.

If steps are fully or partially extended, land as soon as practical.

3-7-C. ELECTRICALSYSTEM FAILURE

Pull EMER INFLATION handle upon water contact to inflate floats.

3-7-D. ELECTRICALPOWER FAILUREIN STEP

STEP EXTND light will be inoperative and visual verification of step in stow shall be made before floats are armed.

3-13. INADVERTENTFLOAT INFLATIONIN FLIGHT

In event of unintended inflation, reduce airspeed, adjust rate of descent to 200 fpm or less, and land as soon as possible. Avoid landing on terrain which could damage float bags.

3-14. EMERGENCYDITCHING WARNING

CREW AND PASSENGERDOORS SHALL REMAINCLOSED DURING DITCHING.

FLOATSswitch - ARMED; check FLOATS caution light Illuminated. Establish an

FAA APPROVED

autorotative

for all GW.

glide speed of 65 to 70 knots

WARNING

SEVERE NOSE UP PITCHING WILL OCCUR IF EMERGENCY FLOATS ARE INFLATED IN FORWARDFLIGHTOR DESCENT.

At 100 feet altitude, execute a moderate cyclic flare to reduce airspeed. Adjust collective and cyclic sufficiently to touchdown in a nose up attitude with forward speed and rate of descent as low as possible. If emergency floats do not inflate immediately upon water contact, pull EMER INFLATIONhandle.

After water touchdown, complete shutdown, check for damage, and determine if helicopter should be abandoned. NOTE

Takeoff after emergency landing with floats inflated is prohibited.

3-15. EGRESSAFTER DITCHING

Ensure rotor is stopped prior to egress.

Crew egress should be accomplished by jettisoning crew doors.

Passengeregress should be accomplished by pushing out passenger door windows at corners marked EMERGENCY EXIT PUSH HERE.

T Section4 RIN P N O R No I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-55.4

change from basic manual.

T N I MODELS R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-56.3 & 56.4


PT6T-3DF ENGINE (30 MINUTE OEI RATING) 412-706-051

S/N 36020 -

36086

S/N 36087 AND SUBSEQUENT CERTIFIED 08 AUGUST1996

This supplement shall be attached to Model 412 or 412EP Flight Manual (BHT-412-FM-3or BHT-412-FM-4) when PT6T-3DFEngine, 30 minute OEI rating kit has been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1996 INC. BELL* HELICOPTER TEXTRON INC. ANDBELLHELICOPTER RESERVED ALLRIGHTS

POSTOFFICE BOX48.

T FORT WORTH.

JGUST 1996


NOTICEPAGE



LOG OF REVISIONS


0.............

08 AUG 96

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP .............................. A- B ..................................

0 0

PAGE

REVISION NO.

i/il ....................................... 1-18 ..................................

0 0

19/20 ....................................

0

NOTE

Revised text is Indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.


FAA APPROVED



0.............

08 AUG 96

APPROVED:

MANAGER



GENERAL INFORMATION

The PT6T-3DEengine has been redesignated as PT6T-3DFto offer an Increased 30 Minute OEI rating. This 30 Minute OEI rating represents improved OEI capabilities while maximum continuous OEI performance Is reduced from that presented In the Basic Flight Manual.

T N I Section1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-56.3 & 56.4

FAA APPROVED

1-1. INTRODUCTION

PT6T-3DF engine offers an increased 30 Minute OEI rating.

Maximum Continuous and Takeoff OEI Operation 30 Minute OEI

1-11. AMBIENTTEMPERA-

21/2 Minute OEI

103.2% 103.2to 109.2% 106.8% 109.2%

TURES

The maximum sea level ambient air temperature for operation is 51.7°C (+125°F).

1-13. POWER PLANT

Pratt and Whitney Aircraft of Canada, Ltd. PT6T-3DF. NOTE

Operation in an OEI range is intended for emergency use only when one engine becomes Inoperative due to an actual malfunction.

Anytime an engine is operated in an OEI range, an entry shall be made in the helicopter logbook detailing the extent of operation in excess of twin engine takeoff power limits. This does not apply to approved ITT limits for starting or for power assurance checks that require torque settings between 59 and 73.2%.

1-13-A. GAS PRODUCERRPM

Continuous Operation

61 to 103.2%

1-13-C. INTERTURBINE TEMPERATURE Continuous Operation Maximum Continuous Maximum Transient (5 seconds above 810°C) OEI Operation 30 Minute OEI 2'1/2 Minute OEI Maximum for Starting (2 seconds above 960°C)

300 to 810° C 810° C

940°C 810 to 940°C 885°C 940°C 1090°C

1-13-D. ENGINETORQUE

Continuous Operation Maximum Continuous OEI Operation 30 Minute OEI 21/2 Minute OEI

5 to 59% 59% 59 to 81% 77% 81%


Refer to Figure 1-1 for Instrument range markings and Figure 1-2 for placards and decals. 1



FAA APPROVED

TWIN & MAX CONT OEI 103.2% 2.5 MINOEI 109.2% 30 MINOEI 106.8%

412FS56.4 -1-2


T Section2 N I R P N O R I O T F C Section3 U T O OD N R E P E E R R Section C 4 S N G O N I R W 0 O E I F V NLY O BHT-412-FMS-56.3 & 56.4

FAA APPROVED



4-1. INTRODUCTION

Performance data presented herein are derived from engine manufacturer's specification power for PT6T-3DF engine less installation losses.


Power Assurance Check charts (figure 41) are provided to determine If the engines can produce Installed specification power. A power assurance check should be performed daily. Additional checks should be made If unusual operating conditions or Indications arise. The hover check Is

performed prior to takeoff, and in-flight check is provided for periodic in-flight monitoring of engine performance. Either power assurance check method may be selected at the discretion of the pilot. It is the pilots responsibility to accomplish the procedure safely, considering passenger load, terrain being overflown, and the qualifications of persons on board to assist in watching for other air traffic and to record power check data. If either engine does not meet the requirements of the hover or the In-flight power assurance check, published performance may not be achievable. The cause of engine power loss, or excessive ITT should be determined as soon as practical. Refer to Engine Maintenance Manual.



Refer to figure 4-2 for increased Single engine rate of climb - 30 minute power.

BHT-412-FMS-56.3& 56.4

Refer to figure 4-3 for for Single engine rate of climb - maximum continuous power.


FAA APPROVED

MODEL412 POWERASSURANCECHECK- HOVER PT6T-3DF ENGINE

HEATER/ECU- OFF.


THROTTLES: TEST ENGINE- FULL OPEN,FRICTIONED. OTHERENGINE- FLIGHTIDLE.


NIIRPM -97%.

COLLECTIVEPITCH - INCREASEUNTIL LIGHT L ON SKIDSOR HOVERING.DO NOT EXCEED810 ITT, 103.1% NI RPM, OR 73.2% TORQUE.

IF EITHER ENGINE EXCEEDS ALLOWABLE ITT AFTER STABILIZING FOUR MINUTES. PUBLISHEDPERFORMANCE MAY NOT BEACHIEVABLE.CAUSE SHOULDBEDETERMINED AS SOON AS PRACTICAL.

STABILIZEPOWER ONE MINUTE, THEN RECORDPRESSURE ALTITUDE,OAT,ENGINETORQUE AND ITT. ENTERCHART AT INDICATEDENGINETOIROUE,MOVE UP TO INTERSECTPRESSUREALTITUDE,PROC EED TO THE RIGHT TO INTERSECTOUTSIDEAIR TEMPERATU IRE, THENMOVE UP TO READ VALUEFORMAXIMUM ALLOWA ABLEITT.

MAXIMUM ALLOWABLE ITT - °C


412-FMS-56.3 & 56.4-4-1.1

Figure 4-1.



FAA APPROVED

MODEL 412 POWERASSURANCE CHECK- IN-FLIGHT PT6T-3DF ENGINE

ESTABLISH LEVELFLIGHTABOVE 1000 FEETAGL AIRSPEED- 100 KIAS IORVNE, IF LESSI

ENTERCHARTAT INDICATEDENGINETORQUE,MOVE UP TO INTERSECTPRESSURE ALTITUDE.PROCEEDTO THE RIGHT TOINTERSECTOUTSIDEAIR TEMPERATURE, THENMOVE UP TOREAD VALUEFOR MAXIMUM ALLOWABLEITT.

HEATER/ECU- OFF


THROTTLES. TESTENGINE- FULL OPEN.FRICTIONED.


OTHERENGINE- DECREASE SLOWLYUNTILTESTENGINE TORQUEIS WITHIN TESTRANGE.DONOT EXCEED810-C ITT, 103.1% NI RPM,OR 73.2% TORQUE.

NII RPM -97%

IF EITHER ENGINE EXCEEDS ALLOWABLE ITT AFTER STABILIZINGFOUR MINUTES, PUBLISHEDPERFORMANCE MAYNOT BE ACHIEVABLE.CAUSESHOULDBEDETERMINED AS SOONAS PRACTICAL.

STABILIZEPOWER ONE MINUTE IN LEVEL FLIGHT.THEN RECORDPRESSUREALTITUDE,OAT,ENGINETORQUE AND ITT

500

550

MAXIMUM 600

ALLOWABLE ITT 700 650

°C

750

800


412-FMS-56.3 & 56.4-4-1-2

Figure 4-1.


7


FAA APPROVED


70 KIAS GROSSWEIGHT7000LB (3175KG) 30 MINUTE POWER ENGINE RPM 97% HEATER OFF INOPERATIVE ENGINE SECURED GENERATOR 150 AMPS WITH ALL DOORS OPEN OR REMOVED: 1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN


-4

-2

0

2

4

6

8

10

12


Figure 4-2.



BHT-412-FMS-56.3 & 56.4


70 KIAS GROSSWEIGHT8000LB (3629 KG) 30 MINUTE POWER HEATER OFF ENGINE RPM 97% INOPERATIVE ENGINE SECURED GENERATOR 150 AMPS WITH ALL DOORS OPEN OR REMOVED: 1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN

RATE OF CUMB - METERS / SECOND

-4

-2

0

2

4

6

8

10

12

18000 16000


9


FAA APPROVED


30 MINUTE POWER GROSSWEIGHT9000LB(4082KG) ENGINE RPM 97% HEATER OFF GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED WITH ALL DOORS OPEN OR REMOVED: 1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN

RATE OF CUMB - METERS / SECOND

-4

-2

0

2

4

6

8

10

12

20000

18000 16000

RATE OF CUMB - FEET/MINUTE

Figure 4-2.



FAA APPROVED


70 KIAS GROSSWEIGHT10,000LB (4536 KG) 30 MINUTE POWER HEATER OFF ENGINE RPM 97% INOPERATIVE ENGINE SECURED GENERATOR 150 AMPS WITH ALL DOORS OPEN OR REMOVED: 1. CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT/MIN


-4

-2

0

2

4

6

8

10

12

18000 16000


Figure 4-2.


11



T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB

70 KIAS GROSSWEIGHT11,900LB (5398 KG) 30 MINUTE POWER HEATER OFF ENGINE RPM 97% INOPERATIVE ENGINE SECURED GENERATOR 150 AMPS WITHALL DOORSOPENORREMOVED:1. CLIMBSPEEDIS 60KIAS 275FT/MIN 2. RATEOF CLIMBWILL DECREASE RATE OF CLIMB - METERS / SECOND

-4

Figure 4-2.

-2

0

2

4

6

8

10

12



FAA APPROVED


MAXIMUMCONTINUOUS POWER ENG - 97% RPM (N2) GENERATOR 150 AMPS

70 KIAS HEATER OFF INOPERATIVE ENGINE SECURED WITHALL DOORS OPEN OR REMOVED I CLIMB SPEED IS 60 KIAS 2. RATE OF CLIMBWILLDECREASE 275 FT/MIN.

20.000

18.000

16.000

Figure 4-3.




BHT-412-FMS-56.3 & 56.4

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 8000 LB (3629 kg) 70 KIAS MAXIMUMCONTINUOUS POWER HEATER OFF ENG - 97% RPM (N21 GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED WITH ALL DOORSOPEN OR REMOVED 1 CLIMB SPEED IS 60 KIAS 2. RATEOF CLIMBWILLDECREASE 275 FT/MIN

20.000

18.0

Figure 4-3.



15


FAA APPROVED

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 9000 LB (4082 kg) MAXIMUM CONTINUOUSPOWER 70 KIAS ENG- 97% RPM (N2) HEATER OFF GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED WITH ALL DOORS OPENOR REMOVED 1 CUMB SPEEDIS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT /MIN

20.000

Figure 4-3.




BHT-412-FMS-56.3 & 56.4

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 10.000 LB (4536 kg) 70 KIAS MAXIMUMCONTINUOUS POWER HEATER OFF ENG - 97% RPM (N21 SECURED ENGINE INOPERATIVE GENERATOR150 AMPS

WITH ALL DOORS OPENOR REMOVED 1 CLIMB SPEED IS 60 KIAS 2. RATEOF CLIMB WILL DECREASE 275 FT /MIN

Figure 4-3.



17


BHT-412-FMS-56.3 & 56.4

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 1 1.000 LB(4990 kg)

MAXIMUMCONTINUOUSPOWER ENG- 97% RPM(N2) 150 AMPS GENERATOR

WITHALL DOORS OPEN OR REMOVED.


CLIMBSPEED IS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT /MIN

20.000

18.000

16.000

Figure 4-3.

18




FAA APPROVED


MAXIMUMCONTINUOUS POWER ENG - 97% RPM IN2) GENERATOR150 AMPS


WITHALL DOORS OPEN OR REMOVED 1. CLIMBSPEED IS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT /MIN

Figure

4-3.

Single

engine

rate of climb

- maximum

continuous

power

(Sheet

6 of 6)

19/20

T N I * MODEL R P N ROTORCRAFT O R I FLIGHT MANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-57.4

SUPPLEMENT

UNITED KINGDOMREGISTERED HELICOPTERS 412-705-016 AND 412-705-018

CAACERTIFIED 1996 11 NOVEMBER

This supplement shall be attached to Model 412EP Flight Manual when UNITED KINGDOM REGISTERED HELICOPTERS kit has been installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1996 INC. BELL* HELICOPTER ANDBELLHELICOPTER TEXTRON INC A DIVISION OFTEXTRON CANADA LTD ALLRIGHTSRESERVED

Inc of Textron A Subsidiary POST OFFICE BOX 482 · FORT WORTH.TEXAS 76101

11 NOVEMBER 1996


NOTICEPAGE


NP


LOGOFREVISIONS


0 ..............

11 NOV 96

LOGOF PAGES

PAGE

Title ..................................... NP ......................................

REVISION NO. 0 0

PAGE

B .................................. A i/ii ....................................... 1 - 52 ..................................

REVISION NO. 0 0 0

NOTE



CAA APPROVED



0 ..............

APPROVED:

CIVIL AVIATION AUTHORITY SAFETY REGULATIONGROUP AVIATION HOUSE SOUTHAREA GATWICKAIRPORT GATWICK WEST SUSSEX RH6 OYR

11 NOV 96

T N I 1 R Section P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-57.4

CAA APPROVED

1-2. BASIS OF CERTIFICATION

1-9. MANEUVERING

This helicopter is certified under BCAR Group A.

1-9-B. CLIMB AND DESCENT


is helicopter The basic configured approved for fifteen-place seating and is certified for operation under non-icing conditions.

The IMC configured helicopter is certified for operation during day or night non-icing conditions.

Refer to Flight Manual Supplement BHT412-FMS-CAA-10.4 for additional limitations procedures and performance data for GROUPA operations.

1-5. CONFIGURATION

EQUIPMENT 1-5-A. REQUIRED

Appropriate equipment as required by BCAR and AND shall be installed.

OPEN/REMOVED 1-5-C. DOORS

For overwater operations with emergency floats installed all doors shall be installed.

1-7. AIRSPEED The Vno is VNE minus 10 knots (Figure 11).

Maximum IMC rate of climb or descent is 1000feet per minute. Maximum degrees.

IMC approach

slope is 5

1-13. POWER PLANT

Instrument markings (Figure 1-2) NOTE

in maximum Operation OEI or maximum continuous power range is contingency intended for emergency use only, when one engine becomes due to an actual inoperative malfunction. Routine operation in OEI or maximum continuous power maximum contingency range can affect engine warranty and service life. NOTE

Operation in excess of twin engine takeoff power limits is permitted only for maintenance purposes in accordance with the Pratt & Whitney Maintenance Manual or Annual Airworthiness Flight test in accordance with the Test CAA Flight approved Schedule.



1-13-A. GAS PRODUCER RPM (N1) 1-13-A-1.

ONE ENGINE INOPERATIVE (OEI)

Continuous

1037%

Maximum Contingency (2½ minute) Range Maximum Contingency

103.7%to 109.2% 109.2%

1-13-B. POWER TURBINE RPM (N1.) Maximum transient (2 seconds maximum)

101.0%

Maximum for Starting (2 sec. max. above 960 °C)

1090°C

1-13-C-2. ONE ENGINE

INOPERATIVE (OEI)

1-13-C-2-a. Gage PN 212-075-067-115 Maximum Continuous 820 °C Maximum Contingency 820 to 925 °C (2'/2 minute) Range Maximum Contingency 925 °C

NOTE Either ITT gage shall be installed In pairs.

1-13-C. INTERTURBINE TEMPERATURE (ITT)

1-13-C-2-b. Gage PN 212-075-067-119 Maximum Continuous

820 °C

1-13-C-1. TWIN ENGINE

Maximum Contingency (2½ minute) Range

820 to 940 °C

Maximum Contingency

940 °C

OPERATION

1-13-C-1-a. Gage PN 212-075-067-115 Continuous

Maximum Continuous

300 to 810 °C 810 °C

Maximum Transient (5 sec. max. above 810 °C)

925 °C

Maximum for Starting (2 sec. max. above 960 °C)

1090 °C

NOTE Either ITT gage shall be installed in pairs.

1-13-C-1-b. Gage PN 212-075-067-119

Continuous

300 to 810 °C

Maximum Continuous

810 °C

Maximum Transient (5 sec. max. above 810 °C)

940 °C

1-13-D. ENGINE TORQUE 1-13-D-1. ONE ENGINE

INOPERATIVE (OEI)

Continuous

5 to 73.2%

Maximum Continuous

73.2%

Maximum Contingency (2½ minute)

73.2 to 81%

Maximum Contingency

81%

1-15.

1-15-D.

ROTOR

ROTOR RPM (NR) -


OEI

97 to 100%


CAA APPROVED

TWIN

103.1% CONTINUOUS

MAX CONTINGENCYOEI 109.2%

OEI 103.7%

412FSCAA-57.4-1-1


3


T N I R N P O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O CAA APPROVED

n

T Section2 N I R P N O R I O T F C Section3 U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-CAA-57.4

CAA APPROVED


3-3.

LOAD position, the battery will provide approximately 75.9 minutes of electrical power.

ENGINE

3-3-A. SINGLEENGINEFAILURE NOTE

Minimum allowable rotorspeed at maximum gross weight may droop to 76% NRduring OEI landing.

3-7. ELECTRICALSYSTEM 3-7-A. DC POWERFAILURE NOTE

A fully charged battery will provide sufficient electrical power for approximately 22.6 minutes under normal conditions. With emergency load switch in EMERG

3-9. AUTOMATICFLIGHT CONTROLSSYSTEM 3-9-C. AFCS HARDOVEROR ABNORMALCONTROL DISTURBANCE WARNING

IF AP1 OR AP2 FAILS OR IS DISENGAGED, REDUCE AIRSPEED TO 115 KIAS OR LESS AND REMAINDER OF FLIGHT SHALL BE EXECUTEDHANDSON.

T Section4 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O CAA APPROVED

4-5. HOVERCEILING

4-5-A. HOVERCEILING- IGE

Adequate cyclic and directional control are available at the gross weights allowed by the Hover Ceiling IGE charts in relative winds up to 35 knots from any direction at or below 3000 feet H,. Improved control margins will be achieved by avoiding winds in the critical relative wind azimuth areas (figure 4-1). The Hover Celling In Ground Effect charts (figure 4-2) provide the maximum allowable gross weights for hovering IGE at all pressure altitude and outside air temperature conditions with heater on or off. Conversely, the hover ceiling altitude can be determined for any given gross weight.

4-5-B. HOVERCEILING- OGE

The Hover Ceiling Out of Ground Effect charts (figure 4-3) provide maximum weights for hovering OGE at all pressure altitude and outside air temperature conditions with heater on or off. Data presented is valid up to 3000 ft HD. CAUTION

OGE HOVER OPERATION MAY RESULT IN VIOLATION OF HEIGHT-VELOCITYLIMITATIONS.

Some of the OGE hover ceiling charts are divided Into two areas as follows:


AREA A (unshaded area) as shown on the hover ceiling charts presents hover performance for which satisfactory cyclic and directional control have been demonstrated in relative winds of 35 knots from any direction at or below 3000 ft HD. Improved control margins will be achieved by avoiding winds In the critical relative wind azimuth areas. AREA B (shaded area) as shown on hover ceiling charts presents additional hover performance which can be achieved In calm winds or winds outside the critical relative wind azimuth areas. NOTE

Tail rotor or cyclic control margin may preclude operation in AREA B of the hover ceiling charts when the relative wind Is In the respective critical wind azimuth area.

4-6. TAKEOFF DISTANCE

The Takeoff Distance charts (figure 4-4) provide takeoff distance required to clear a 50 foot (15 meter) obstacle in a zero wind condition, using a takeoff flight path which will avoid the critical areas of the HeightVelocity diagram (Section 1). Takeoff Is Initiated from a hover at 4 feet (1.2 meters) skid height with climbout speed of 45 knots. NOTE

Downwind takeoffs are not recommended because the published takeoff distance performance cannot be achieved.



CAA APPROVED

weight combinations with heater off at maximum continuous power.

NOTE

NOTE

When climbing at 100% mast torque and OVER TORQUE light illuminates, reduce power slightly.

Published single engine performance is intended for emergency use only when one engine becomes inoperative due to an actual malfunction. Routine operation in maximum contingency (2 1/2minute) OEI or maximum contingency OEI range can affect engine service life.

4-7-A. TWINENGINERATE OF CLIMB

The Twin Engine Rate of Climb charts (figure 4-5) provide the rates of climb that can be obtained at all outside air temperatures/pressures altitudes/gross weight combinations with heater on or off at maximum continuous power and takeoff power.

NOTE

OEI operations is however permitted for maintenance, OEI training, and airworthiness flight test purposes.

4-9.

NOTE

All rate of climb data are based on changes in true altitude (pressure altitude corrected for nonstandard temperature).

4-7-B. SINGLEENGINERATE OF CLIMB

The Single Engine Rate of Climb charts (figure 4-6) provide the rates of climb that can be obtained at all outside air temperature/pressure altitudes/gross

LANDING DISTANCE

The Single Engine Landing Distance chart (figure 4-7) provides the landing distances required to clear a 50 foot (15 meter) obstacle for all outside air temperatures, pressure altitudes, and gross weights. Landing distances are based on an approach condition of 45 KIAS and 500 feet per minute rate of descent and zero wind.


270


NOTE: 1. Pedal critical wind azimuth-hovering with the relative wind within SEENOTE these azimuth angles can result in inability to maintain heading due to large left pedal requirements for certain wind velocities. 2. Longitudinal cyclic critical wind azimuth -

°

95

aft

cyclic

limited longitudinal hardover.

may

be

with AFCS

SEENOTE 2

OGE CRITICAL RELATIVE WIND AZIMUTH

SEENOTE 1

270

SEENOTE 2

IGE CRITICAL RELATIVE WIND AZIMUTH

Figure 4-1.

412-FS-CAA-57.4

4-1

Critical relative wind azimuths

9


CAA APPROVED


POWER: SEE NOTE BELOW ENG - 100% RPM IN2) GENERATOR 150 AMPS (EA.)

SKID HEIGHT 4 FEET HEATER ON OR OFF ° -40 TO 52°C

NOTE THESE IGE HOVER CEILINGS ARE BASED ON DENSITY ALTITUDE LIMITS FOR TAKEOFF AND LANDING THIS HELICOPTER CAN BE HOVERED IGE AT THE INDICATED GROSS WEIGHTS WITH LESS THAN TAKEOFF POWER FOR TEMPER. ATURES BELOW 48°C MAXIMUM

GROSS

WEIGHT

LIMIT

MAX.

OAT HEATER ON (21°C)

11.9

-40-30-20-10

0

10

20

OAT - °C

Figure 4-2.

50

60

8

3.5

9

10

11


Hover ceiling in ground effect (Sheet 1 of 2)

12 LB X 1000

5.4

kg X 1000



HOVER CEILING

IN GROUND

EFFECT

SKID HEIGHT 4 FEET HEATERON OR OFF ° -40 TO 52°C

MAXIMUMCONTINUOUSPOWER ENG- 100% RPM (N2) GENERATOR 150 AMPS(EA.)


MAX GROSS WEIGHT LIMIT OAT OAT0°C C

MAX OAT

AX OAT

HEATER ON (21°C)

-40-30-20

-10 0

10 20 30 40 5060

8

OAT -°C

10

11

12 LBX 1000


5.4 kg X 1000

.

3.5

Figure 4-2.

9

Hover ceiling in ground effect (Sheet 2 of 2)

11




TAKEOFFPOWER ENG- 100% RPM(N2) GENERATOR 150 AMPS (EA.)

SKID HEIGHT 60 FT. HEATEROFF 0 TO 52°C

CAUTION: OGEHOVEROPERATION MAY RESULTIN VIOLATIONOF H-VLIMITATIONS.

AREA A

LIMIT

60

8

3.5

Figure 4-3.

9

10

11


12 LB X 1000

5.4 kg X 1000





SKID HEIGHT 60 FT. HEATEROFF -40 TO O°C

TAKEOFFPOWER ENG - 100% RPM (N2) GENERATOR 150 AMPS (EA.)

CAUTION: OGEHOVEROPERATIONMAY RESULTIN VIOLATIONOF H-VLIMITATIONS. 14,000 FT. DEN. ALT. LIMIT

-30-20-10 OAT°C

0

8

3.5

Figure 4-3.

9

4.0

10

4.5

GROSS WEIGHT

11

5.0

12 LBX 1000

5.4

kg X 1000


13


CAA APPROVED


TAKEOFFPOWER ENG- 100% RPM (N2) GENERATOR 150 AMPS (EA.)


CAUTION: OGEHOVEROPERATION MAY RESULTIN VIOLATIONOFH-V LIMITATIONS.

TORQUE LIMIT

O 10 20 OAT -C

8

3.5

Figure 4-3.

9

10

11

12 LB X 1000

.


5.4


kg X 1000




SKID HEIGHT 60 FT. HEATER ON ° -40 TO 0 C

TAKEOFFPOWER ENG- 100% RPM (N2) GENERATOR 150 AMPS(EA.)

CAUTION: OGEHOVEROPERATIONMAY RESULTIN VIOLATIONOF H-VLIMITATIONS. LIMIT

TORQUE LIMIT

8

OAT -

3.5

Figure 4-3.

9

10

11

12 LB X 1000

°C


5.4 kg X 1000


15


CAA APPROVED


MAXIMUM CONTINUOUSPOWER ENG- 100% RPM (N2) GENERATOR 150 AMPS(EA.)


CAUTION: OGEHOVEROPERATIONMAY RESULTIN VIOLATIONOFH-V LIMITATIONS.

OAT- °C

Figure 4-3.

3.5

.

. . . 4.0 4.5 5.0 5.26 GROSS WEIGHT


kg X 1000


CAA APPROVED


SKID HEIGHT 60 FT. HEATER OFF ° -40 TO 0 C

MAXIMUMCONTINUOUSPOWER ENG- 100% RPM(N2) GENERATOR 150 AMPS(EA.)

CAUTION: OGEHOVEROPERATIONMAY RESULTIN VIOLATIONOF H-VLIMITATIONS. 14.000 FT DEN. ALT LIMIT

LIMIT _

-40 -30 -20 -10 OAT°C

0

8

3.5

Figure 4-3.

9

10

11

12 LB X 1000

5.26

4.0 4.5 5.0 5.26 GROSS WEIGHT

kg

kg X 1000


17


CAA APPROVED


MAXIMUM CONTINUOUSPOWER ENG - 100% RPM (N2) GENERATOR 150 AMPS (EA.)

SKID HEIGHT 60 FT. HEATERON 0 TO 20°C

CAUTION: OGEHOVEROPERATION MAY RESULTINVIOLATIONOF H-V LIMITATIONS.

0

10 20

8

3.5

Figure 4-3.

9

10

11


12 LB X 1000

5.26


kg X 1000


CAA APPROVED


SKID HEIGHT 60 FT.

MAXIMUMCONTINUOUSPOWER

CAUTION OGEHOVEROPERATIONMAY RESULTIN VIOLATIONOF H-VLIMITATIONS.

LIMIT

8

3.5

9

4.0

10

4.5

11

12 LBX1000

5.0 5.26

GROSS WEIGHT

Figure 4-3.


kg X 1000


CAA APPROVED

TAKEOFF DISTANCE OVER 50 FOOT OBSTACLE

HOVER POWER + 15% TORQUE ENG - 100% RPM IN2) GENERATOR 150 AMPS IEA.)

INITIATED FROM 4 FT SKID HEIGHT VTOCS= 45 KIAS HEATER ON OR OFF

14,000 FT. DEN ALT. LIMIT

-MAX

OAT

AT

MAX OAT HEATER

ON

(21-C)

i0-40-20

0

20 40

OAT - 'C

60 400

600

800

1000

TAKEOFF DISTANCE

1200

-

ENGLISH

Figure 4-4.

Takeoff distance (Sheet 1 of 2)

FT.

1400



HOVERPOWER+ 15% TORQUE ENG - 100% RPM(N2) GENERATOR 150 AMPS(EA.)

Figure 4-4.

INITIATED FROM 1 2 meter SKID HEIGHT VTOCS= 45 KIAS

HEATERON OR OFF

Takeoff distance (Sheet 2 of 2)


CAA APPROVED


TAKEOFFPOWER

70 KIAS

GENERATOR 150 AMPS(EA.)

2 RATEOF CLIMB WILL DECREASE



Figure Figure 4-5. 4-5.

22

Twin Twin engine engine rate rate of of climb climb (Sheet (Sheet 11 of of 24) 24)


CAA APPROVED


GROSS WEIGHT7000 LB (3175 kg)

70 KIAS HEATERON

TAKEOFF POWER ENG- 100% RPM(N2) GENERATOR 150 AMPS (EA.)

WITHALL DOORSOPENOR REMOVED

CLIMB SPEEDIS 60 KIAS 2 RATEOF CLIMB WILL DECREASE 275 FT /MIN.

20.000

-6000

18.000

OAT LIMIT

16,000

AT -20°C 30C -40 C

-5500

-5000 -4500

-4000

I

12.000

-3500

10.000

- 3000 -2500

8,000

- 2000

6,000

\

-1500

4,000

-1000

2,000

0

-500

400


(2.0)

(4.0) (6. (80) .0) (10.0) (12.0) RATE OF CLIMB - (METERS/SECOND)

Figure 4-5.

2800

(14.0)

3200

(16.0)


23



HEATEROFF

ENG- 100% RPMIN21

275 FT. MIN

20.000

-6000

-1500

0

RATE OF CLIMB -

FEET/MINUTE

0)

RATE OF CLIMB -

Figure 4-5.

(METERS/SECOND)





70 KIAS HEATERON

TAKEOFFPOWER ENG- 100% RPM (N21 150 AMPS (EA.) GENERATOR


1 CLIMB SPEED IS 60 KIAS 2 RATE OF CLIMB WILL DECREASE 275 FT /MIN


Figure 4-5.


25


CAA APPROVED


TAKEOFFPOWER 70 KIAS HEATER OFF ENG- 100% RPM 9N2) 150 AMPS(EA.) GENERATOR WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATE OF CLIMBWILL DECREASE 275 FT /MIN

20.000

6000 6000

18000

-5500

16000

OAT

5000

-20 C 30°C

4500

14,000

-4000

-4000

12,000

10,000

3000

-3000

-2500

8.000

OAT LIMIT

0

400


(0)

(2.0)


Figurer 4-5.

Twin

engine

rate

of Climb

(sheet

2800

(14.0)

5 Ot 24)

3200

(16.0)



GROSS WEIGHT

9000

LB (4082

kg)

70 KIAS

TAKEOFF POWER

GENERATOR 150AMPS IEA.I



OAT

-5500

-5000

-4500

cc

L

L

-2000

K

-1500 -1000 -500

Figure 4-5.


27




GROSS WEIGHT 10.000 LB (4536 kg) TAKEOFFPOWER ENG - 100% RPM (N2) 150 AMPS (EA.) GENERATOR

70 KIAS HEATEROFF

WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATE OF CLIMBWILL DECREASE 275 FT /MIN

20,000

18,000

16,000

14,000

OAT LIMIT




TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10,000 LB 14536 kg)

70KIAS HEATER ON

TAKEOFF POWER ENG - 100% RPM (N2) GENERATOR 150 AMPS (EA.)



20.000

-6000 -5500 -5000 -4500

000

-5 00

0

0

400


(0)

(2.0)


Figure 4-5.

2800

14.0)

3200

(16.0)



CAA APPROVED



30




GROSS WEIGHT 11,000 LB (4990 kg) TAKEOFFPOWER ENG - 100% RPM (N2) 150 AMPS(EA.) GENERATOR

70 KIAS HEATER ON

WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATEOF CLIMB WILL DECREASE 275 FT /MIN

20,000

I

(2.0)

(6.0) (8.0) (10.0) (12.0) (14.0) (16.0) (4.0) RATE OF CLIMB - (METERS/SECOND)

Figure 4-5.


31


32






WITHALL DOORSOPENOR REMOVED 1 CLIMBSPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT /MIN

20,000

OAT

Figure 4-5.


33


CAA APPROVED

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 7000 LB (3175 kg) MAXIMUM CONTINUOUSPOWER ENG - 100% RPM (N2) 150 AMPS(EA.) GENERATOR WITH ALL DOORS OPEN OR REMOVED

70 KIAS HEATER OFF


20.000

-6000 -5500

18,000

-5000

16,000

-

14,000

4.000

2,000

0

Figure 4-5.

34



CAA APPROVED


MAXIMUMCONTINUOUSPOWER ENG- 100% RPM (N2) 150 AMPS(EA.) GENERATOR


70 KIAS HEATER ON


20,000

-6000

18,000

-5500

OAT

-5000

16.000

-4500

c

14.000

I

-1500

Figure 4-5.


35





MAXIMUM CONTINUOUSPOWER ENG- 100% RPM(N2) 150 AMPS(EA.) GENERATOR

70 KIAS HEATEROFF

WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT /MIN

20,000

-6000

18,000

-5500

-5000

16,000

-4500

14,000

LU

-4000

12.000

-3500

10,000

8.000

6.000

4,000

1000

2,000

500

0


400

2800

(0)


Figure 4-5.


3200


CAA APPROVED

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 8000 LB 13629 kg) MAXIMUMCONTINUOUSPOWER ENG- 100% RPM (N21 150 AMPS (EA.) GENERATOR


70 KIAS HEATER ON


20,000

18,000

16,000

1500

1000

500

0


400

2800

3200


Figure 4-5.


37




70 KIAS MAXIMUM CONTINUOUSPOWER ENG - 100% RPM(N21 HEATER OFF GENERATOR 150 AMPS(EA.) WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATE OF CLIMBWILL DECREASE 275 FT /MIN

0

200

(0)

(1.0)


(2.0)

(3.0)

(4.0)

(5.0)

(6.0)

1400

(7.0)

1600

(8.0)


Figure 4-5.



CAA APPROVED

TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 9000 LB (4082 kg) MAXIMUM CONTINUOUSPOWER ENG - 100% RPM (N2) 150 AMPS IEA.) GENERATOR

70 KIAS HEATER ON

WITH ALL DOORSOPENOR REMOVED 1. CLIMB SPEEDIS 60 KIAS 2. RATE OF CLIMB WILL DECREASE 275 FT /MIN

20,00

0

0


200

1400


Figure 4-5.


1600




GROSS WEIGHT 10,000 LB (4636 kg) MAXIMUM CONTINUOUSPOWER ENG- 100% RPM(N2) GENERATOR 150 AMPS (EA.)

70 KIAS HEATER OFF

WITHALL DOORSOPENOR REMOVED 1. CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT /MIN

20,000

1000

500


Figure 4-5.

40




TWIN ENGINE RATE OF CLIMB GROSS WEIGHT 10.000 LB (4536 kg)

MAXIMUMCONTINUOUSPOWER ENG- 100% RPM (N2) 150 AMPS(EA.) GENERATOR

70 KIAS HEATER ON

WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2. RATEOF CLIMBWILL DECREASE 275 FT /MIN.

0

200

(0)

(1.0)



Figure 4-5.

1400

(7.0)

1600

(8.0)


41


CAA APPROVED

TWIN ENGINE RATE OF CLIMB. GROSS WEIGHT 11,000 LB (4990 kg)

MAXIMUMCONTINUOUSPOWER ENG- 100% RPM IN2) GENERATOR 150 AMPS(EA.)

70 KIAS HEATER OFF



20,000

18,000

OAT

- 3000

1000

0

200

(0)

(1.0)

Figure

42



4-5.

Twin

engine

rate of climb

(Sheet

1400

(7.01

21 of 24)

1600

(8.0)

I




GROSS WEIGHT

11.000

LB (4990

kg)

MAXIMUM CONTINUOUS POWER ENG - 100% RPM IN2) GENERATOR 150 AMPS (EA.)


70 KIAS HEATER ON


20,000

18,000

I LU

RATE OF CLIMB -

(0)

(1.0)

FEET/MINUTE


Figure 4-5.

(7.0)

(8.0)


43




GROSS WEIGHT 11.900 LB (5398 kg) MAXIMUMCONTINUOUSPOWER ENG- 100% RPMIN2) GENERATOR 150 AMPS (EA.I

70 KIAS HEATEROFF


(O)

(110)

(2.0) (3.0) (4.0) (5.0) (6.0) (7.0) RATE OF CLIMB - (METERS/SECOND)

Figure 4-5.

44

(8.0)



CAA APPROVED


GROSS WEIGHT 11,900 LB (5398 kg) MAXIMUM CONTINUOUSPOWER ENG- 100% RPM (N2) GENERATOR 150 AMPS (EA.)

70 KIAS HEATER ON

WITHALL DOORSOPENOR REMOVED 1 CLIMBSPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT /MIN.

20.000

18,000

1000

500

Figure 4-5.


45


CAA APPROVED


MAXIMUM CONTINUOUSPOWER ENG- 97% RPM (N2) GENERATOR 150 AMPS


WITHALL DOORS OPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT /MIN

20,000

-6000

18,000

-5500 -5000

16,000

-4500

14,000

LU

12,000

1200


Figure 4-6.



CAA APPROVED

SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT

8000

MAXIMUM CONTINUOUS POWER ENG - 97% RPM (N2) GENERATOR 150 AMPS


LB (3629

kg)

INOPERATIVE

70 KIAS HEATER OFF ENGINE SECURED


20,000

-6000 -5500 -5000 -4500

-

-1500

-1000 -500

1200 (6.0)

Figure 4-6.


47

CAA APPROVED


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 9000 LB (4082 kg) 70 KIAS MAXIMUMCONTINUOUSPOWER HEATER OFF ENG - 97% RPM (N21 GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED WITHALL DOORSOPENOR REMOVED 1 CLIMB SPEEDIS 60 KIAS 2 RATEOF CLIMBWILL DECREASE 275 FT /MIN

20.000

-6000 -5500

-5000

-4600

w

800

(4.0)


Figure 4-6.


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O SINGLE ENGINE RATE OF CLIMB

GROSS WEIGHT 10.000 LB (4536 kg) MAXIMUMCONTINUOUSPOWER

70 KIAS


Figure 4-6.


49




GROSS WEIGHT 11,000 LB (4990 kg) MAXIMUM CONTINUOUSPOWER 70 KIAS ENG- 97% RPM(N21 HEATEROFF GENERATOR 150 AMPS INOPERATIVE ENGINE SECURED


20.00

-6000

18.00

-5500

-5000

16.00

-4500

-2000

-1600

-1200

-800

-400

0

400


(2.0)


Figure 4-6.

50




SINGLE ENGINE RATE OF CLIMB GROSS WEIGHT 11.900 LB (5398 kg)

MAXIMUM CONTINUOUS POWER ENG - 97% RPM (N21 GENERATOR 150 AMPS

INOPERATIVE


70 KIAS HEATER OFF ENGINE SECURED


20.000

18,000

16,000

TWIN ENGINE M C P ABSOLUTE CEILING -

14.000

4000

1000

2000

500

-2000

-1600

-1200

RATE OF CLIMB -

(-10.0)

(-8.0)

(-6.0)

-800

-400

0

400

(0)

(2.0)

FEET/MINUTE (-4.0)

(-2.0)


Figure 4-6.


51



SINGLE ENGINE LANDING DISTANCE OVER 50 FT. (15 M) OBSTACLE

21/2 MINUTE OEI POWER AS REQUIRED ENG - 97% RPM (N2) GENERATOR 150 AMPS

Figure 4-7.

Single engine landing distance

T N I R P N ROTORCRAFT O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-58.4


EXTERNAL CLASS D LOADS

ALL MODEL 412/412EP WITH PT6T-3D,-3DE OR -3DF ENGINE CERTIFIED 12 MARCH1997 All operations shall be conducted under FAR Part 133 Rotorcraft External Load Operations, Subpart D Airworthiness Requirements. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1997 BELLa HELICOPTER TEXTRON INC. TEXTRON INC. ANDBELLHELICOPTER LTD. CANADA OFTEXTRON A DIVISION ALLRIGHTS RESERVED

A Subsidiary of TextronInc POST OFFICE BOX 482

· 76101 FORT WORTH TEXAS

12 MARCH 1997


NOTICEPAGE


0


LOG OF REVISIONS


0 .............

12 MAR 97

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP .............................. AB ..................................

0 0

PAGE

i/ii ....................................... 1 - 4 ................................... 5/6 ......................................

REVISION NO. 0 0 0

NOTE



FAA APPROVED



0.............

12 MAR 97

APPROVED:



GENERAL INFORMATION

Class D operations permit hoisting of a person up or down from a hovering helicopter in a non-emergency situation. To accomplish a Class D operation with a Transport Category A Degree of Safety, the following conditions must be met: * * *

No single failure will terminate the maneuver. Power required for establishing OEI hover does not exceed any engine or transmission limits, and maneuver shall be completed within 2.5 minute OEI limitations. Loss of altitude, after any failure and normal pilot recognition, does not exceed 10% of hover height.



Hoist operations shall be conducted under appropriate operating rules for external Class D loads (FAR PART 133).

Hoist operations are limited to day/night visual meteorological conditions (VMC).

Passenger operations during Class D hoist operations are prohibited except for those passengers to be hoisted.

BHT-412-FMS-58.4

Helicopter shall be equipped with PT6T3D,-3DE or -3DF engine and ITT gages (212-075-067-115). NOTE

ITT gage P/N 212-075-67-115shall be installed in pairs

1-5-B. CLASS D OPERATIONS

Hoist system shall be FAA approved.

1-4. FLIGHT CREW

The FAA approved hoist system must continue to operate normally following failure (electrically engine/generator connected to a non-shed bus).

Minimum flight crew shall consist of one pilot approved for type of external operations and one copilot/crewmember who shall act as hoist operator.

Two engine generator load limits must be established such that following failure of the critical engine or generator, remaining generator will remain within flight manual limits during hoist operations.

1-5. CONFIGURATION

Hoist system utilized for carrying Class D loads shall require two distinct crew actions to jettison the load.

1-5-A. REQUIREDEQUIPMENT

acting as hoist Copilot/crewmember operator shall wear an approved safety harness in passenger compartment during all phases of hoist operations. Hoist operator shall wear protective gloves for guiding cable during operations. Hoist operator shall be familiar with hoist operating procedures and limitations. shall have cabin Hoist operator intercommunication with pilot at all times.

H-V diagram shown in limitations section of basic flight manual is not applicable while operating in accordance with this supplement.

All other limitations shall be in compliance with FAR Part 133 where applicable.

T N I Section 2 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W Section 3 O E I F V NLY O BHT-412-FMS-58.4

FAA APPROVED


Gross weight and CG - Compute with and without hoist load.

2-4. INTERIOR AND PRESTART CHECK

2-4-A. HOISTINSTALLATION CHECKS

CAUTION

SEE BASIC FLIGHT MANUAL FOR HOIST PENALTY REGION LIMITATION WITH ASSOCIATED HOIST CG ENVELOPE.

Refer to appropriate Flight Manual Supplement for required checks, limitations, and operating procedures. For Bell Helicopter approved hoist, refer to BHT-412-FMS-7 and -26.

2-7. BEFORE TAKEOFF

Perform power assurance checks, refer to Flight Manual.



FAA APPROVED

DOPERATIONS: CLASS

4-11. OPERATION

4-11-B. HOVER HEIGHT

Prior to commencement of an approach for a Class D operation, three additional factors must be determined. These three factors are:

If intended operation requires vertical load/cable retrieval while hovering, pilot must determine maximum hover height based on hoist/cable retrieval speed. Flight tests have demonstrated maximum hover height should not exceed distance hoist cable travels in 90 seconds during retrieval with intended hoist load.

* Single engine (target) torque required to remain in an OEI hover. * Hover height required for intended operation based upon hoist speed during OEI. * Helicopter gross weight at which an OEI hover can be maintained at 2.5 minute power.

4-11-A. SINGLE ENGINE (TARGET) TORQUE

To determine single engine (target) torque available for OEI hover, enter Class D External Load Gross Weight for Hovering chart (figure 4-1) at pressure altitude, proceed horizontally to OAT, then down to engine (target) torque. NOTE

calculated Engine (target) torque will be 1.5 % or, 20° ITT less than power minimum specification available to ensure that altitude loss with normal reaction time will be less than 10% of hover than A better altitude. specification engine, or a wind component may reduce power required for hover and a reduction of power may be required to maintain level hover.

NOTE

Remaining time at 2.5 minute OEI power is to be used for stabilization and transition to forward flight (Vy).

If a vertical load/cable retrieval is not required and a clear approach/departure corridor is available, then hover height can be at pilot discretion.

4-11-C. GROSS WEIGHT

To determine maximum gross weight which can be sustained in an OEI hover, enter chart (figure 4-1) at pressure altitude, to OAT. Drop proceed horizontally vertically to desired skid height for intended operation. Proceed horizontally to determine maximum gross weight. NOTE

During the CG/gross weight planning phase of a Class D operation, two key factors need to be retained. The first is fuel


weight at the maximum OEI gross weight, and the second is engine (target) torque for the intended operation. If an engine fails during hoist operations at or below that

FAA APPROVED

fuel weight, a collective adjustment to the engine (target) torque will sustain the hover without exceeding any engine limits.


FAA APPROVED

CLASS 'D' EXTERNAL LOAD GROSS WEIGHT FOR HOVERING

MODEL:412/412EP DATE: 28 FEBRUARY1997 DATA: FLIGHTTEST

2.5 MINUTE OEI POWER 96% ROTOR RPM ZERO WIND

ENGINE: PT6T-3D/-3DE/-3DF

FUELGRADE:JP-4/ JP-5/JP-8 FUELDENSITY:6.5/ 6.8/6.8 LB/GAL

8000 7000

12

54 58 62 66 70 74 78 82 ENGINE(TARGET) TORQUE -%

Figure 4-1.

Class D external load gross weight for hovering

5/6

T N I R P N ROTORCRAFT O R I FLIGHT MANUAL O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-59.4

SUPPLEMENT

TRIMBLE 2101 GPS WITH

ELECTRICAL/MECHANICAL DISPLAY FOR VFR OPERATIONS 412-899-510 CERTIFIED 17 JUNE 1997

This supplement shall be attached to Model 412EP Flight Manual when Trimble 2101 Approach Global PositioningSystem has been installed. Informationcontained herein supplements information of basic Flight Manual. For Limitations,Procedures,and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

NOTICE COPYRIGHT COPYRIGHT 1997 INC TEXTRON BELL HELICOPTER ANDBELLHELICOPTER TEXTRON INC. A DIVISION OFTEXTRON CANADA LTD

of Textron Inc A Subsidiary 76101 TEXAS OFFICEBOX482 · FORTWORTH, POST

17JUNE

1997


NOTICEPAGE





LOG OF REVISIONS


0............

17 JUNE 97

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP .............................. A - B .................................. i/ii ....................................... 1- 2 ...................................

REVISION

PAGE

NO.

3/4 ......................................

0 0 0 0

0

MANUFACTURER'S DATA

5/6 ......................................

0

NOTE





0............

17 JUNE 97

APPROVED:

MANAGER

ROTOR RAFT CERTIFICATIONOFFICE FEDERALAVIATION ADMINISTRATION FT. WORTH,TX 76193-0170

DATE:

6-17-97


GENERAL INFORMATION

Global Positioning System (GPS) is a global coverage satellite based navigation system. It provides standard positioning service (SPS) to all equipped users. Trimble 2101Approach is a long range, GPS based RNAV system consisting of a DZUS-rail mounted receiver/ display unit and an antenna. This system provides pilot with navigational guidance and position information in three dimensions: latitude, longitude, and altitude. Every 28 days Trimble makes Jeppesen data base NavData card updates available to Trimble 2101 Approach GPS users.

Trimble 2101 Approach has been designed to meet accuracy specification of VFR enroute oceanic and remote enroute domestic, terminal, and non-precision instrument approach operation within U.S. National Airspace System, North Atlantic Minimum Navigation Performance Specification (MNPS)Airspace and latitudes bounded by 74° North and 60° South using WGS-84 (or NAD 83) coordinate reference datum in accordance with criteria of AC 20-138, AC 91-49, and AC 120-33. Navigation data is based upon use of only global positioning system (GPS) operated by the United States. System accuracy is predicated on system receiving usable signals. NOTE

This installation is approved for VFR only.

T N I Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R 2 EW Section O I F V NLY O BHT-412-FMS-59.4

FAA APPROVED

1-1. INTRODUCTION

System must utilize Software number 0237.

Limitations described In this supplement are in additions to those described in basic flight manual and apply only to model 412EP helicopter equipped with GPS kit installed.


Trimble 2101 Approach GPS Pilot's Guide, (Publication Number 81451) dated November 1996 must be Immediately available to flight crew whenever navigation is predicated on use of system. Software number of Pilot's Guide must match System Code annunciated on System,Code page.

1-3. TYPES OF OPERATION NOTE

This operation VFR only

is approved for

1-6-B. CENTEROF GRAVITY


1-20. INSTRUMENT MARKINGSAND PLACARDS

GPS LIMITEDTO VFRUSE ONLY

Location: Pilot instrument panel; Copilot instrument panel.

2-4. INTERIOR AND PRESTARTCHECK

*Add

2-4-A-1.

2-6. SYSTEMSCHECK

LOWER PEDESTAL

GPS POWER switch -

OFF

2-6-H. GPS CHECK 1

T N I R P N O R I O T F C U T D O Section 3 O N R E P E E R R C S N G O N Section4 I R W O E I F V NLY O BHT-412-FMS-59.4

GPS POWER switch -

FAA APPROVED

follow procedure described in Trimble 2101 Approach GPS Pilot's Guide. This altitude should be within 100 ft of actual geographic altitude of A/C.

ON

During Self-Test -:

Verify GPS annunciators (APR, HLD, WPT and MSG) are illuminated Once Self-Test Is passed -:


Normal operating procedures are outlined In Trimble 2101 Approach GPS Pilot's Guide, (Publication Number 81451) dated November 1996.

Verify proper barometric altitude is displayed on Altitude page; to do so,

3-15. GPS MESSAGES

message indicates GPS malfunction or degradation of capability.

Refer to Trimble 2101 GPS Pilot's Guide. Discontinue VFR GPS operations, If


BHT-412-FMS-59.4

T Section5 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O No change from basic manual.

3/4

T Section1 RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'SDATA

1-17. NAVIGATIONSYSTEM AND INSTRUMENTS

1-17-B. OPTIONALNAVIGATION EQUIPMENT

1-17-B-9. GLOBAL POSITIONING SYSTEM (GPS) Remote switches OBS knob knob.

Pilots HSI course select

There is no provision for HSI analog course input in TRIMBLE 2101 APPROACH. To properly use deviation bar on HSI, pilots should manually set HSI pointer to Desired Track (DTK) on TRIMBLE 2101 APPROACH display, after selecting GPS as

BHT-412-FMS-59.4

source with navigation pushbutton annunciator.

CRS PTR

APR/HLD/WPT/MSG- Approach, Hold, Waypoint and Message annunciator located on instrument panel.

Annunciator will illuminate to alert pilot of TRIMBLE 2101 APPROACHon activation of Approach mode (APR, green), on Holding mode of active waypoint (HLD, white), on turn anticipation (WPT, amber), and of a Message (MSG amber). These four (4) lights will illuminate during system selftest.

CRS PTR/NAV/GPS- Navigation Source located on Switch/Annunciator instrument panel.

Use to select navigation source display on pilots HSI. Press to select desired source.

5/6



SLOPE LANDING OPERATIONS 412-704-012 CERTIFIED 09 JULY 1997

This supplement shall be attached to the 412 Flight Manual (BHT-412-FM-3)or Model 412EP Flight Manual (BHT-412-FM-4)when 412-704-012Slope Landing Kit is installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

Bell Helicopter

COPYRIGHT NOTICE COPYRIGHT 1997 ® TEXTRON INC BELL HELICOPTER INC TEXTRON ANDBELLHELICOPTER CANADA LTD OFTEXTRON A DIVISION ALLRIGHTS RESERVED

A Subsidiary of Textron Inc POSTOFFICE BOX 482 - FORT WORTH, TEXAS 76101

REISSUE -

29 JULY 1997


NOTICEPAGE



LOG OF REVISIONS


0..............09 0 ..............

JUL 97 29 JUL 97

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title -

NP ..............................

0

PAGE

REVISION NO.

A - B .................................. 0 i/ii . .. . . . . .. . . . . . . .. .. . . . . . . . . .. . .. . . . . . .. 0 1/2 ...................................... 0

NOTE




Original ........... Reissue...........0

0..............09 ..............

JUL 97 29 JUL 97

APPROVED:

MANAGER


DATE:


GENERAL INFORMATION

This supplement provides the additional limitations when the Slope Landing Kit 412-704012 is installed. This kit consist of a Main Rotor Mast Assembly with a 5000 hour retirement life.

i/ii


FAA APPROVED

*

1-9. MANEUVERING

touch ground. If stinger touches abort the landing. NOTE

1-9-C. SLOPELANDING

Slope landings are limited to a maximum slope angle of 10 degrees and 8 degrees, figure 1-1.

When set for skids level attitude, the helicopter attitude indicator may be used as a reference for determining the maximum slope angle.

CAUTION

When landing at aft longitudinal center of gravity, tail stinger may

ALL HEADINGS INDICATE THEDIRECTION OF THENOSEOF THEHELICOPTER NOSE UPSLOPE

0

90

180'

NOSE DOWNSLOPE

Figure 1-1. Slope Landing

412-FMS-1-1-1

T N I R P N ROTORCRAFT O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4


CATEGORYA OPERATIONS WITH PT6T-3D SERIES ENGINE INSTALLATIONS S/N 36020 - 36086

S/N 36087 AND SUB CERTIFIED 23 APRIL 1999

This supplement shall be attached to Model 412/412EP Flight Manualwhen operating IAWCategoryA and PT6T3D series engines have beeninstalled. Information contained herein supplementsinformation of Basic Flight Manual.For Limitations, Procedures,and Performance Data not contained in this supplement, consult Basic Flight Manual.

COPYRIGHT NOTICE BELL

Bell

Helicopter

HELICOPTER TEXTRON INC.

A DIVISION OFTEXTRONCANADALTD. ALL RIGHTSRESERVED

23 APRIL 1999 REVISION1 - 2 JUNE 1999


NOTICEPAGE


Manufacturer's Data portion of this supplement is proprietary to Bell Helicopter Textron Inc. Disclosure,reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorizationfrom Bell HelicopterTextron Inc.

Additional copies of this publication may be obtained by contacting: CommercialPublication Distribution Center Bell Helicopter Textron Inc. P. O.Box 482 Fort Worth, Texas76101-0482

NP


LOG OF REVISIONS

0.................... 23 APR 99 Original .. ..................... Revision ...................... 1...................... 2 JUN 99

LOG OF PAGES

REVISION NO.

PAGE

PAGE

MANUFACTURER'S DATA

FLIGHT MANUAL

Title ........................................ ................. 1 NP ..............................................................0 A ................................................................1 B .................................................................0 i- ii...........................................................0 1 - 96 ........................................................0 97/98 ..........................................................0

REVISION NO.

1 99 ............................................................... 100 - 118..................................................0

NOTE


A


FAA APPROVED

LOG OF FAA APPROVED APPROVEDREVISIONS

0..................23 APR 99 Original........................

APPROVED

ROTORCRAFT CERTIFICATION OFFICE FEDERALAVIATIONADMINISTRATION FT.WORTH,TX 76193-0170 B

DATE


BHT-412-FMS-62.3 & 62.4

GENERAL INFORMATION INTRODUCTION

Category A operations have been separated into four parts:

minimum, or use Drop Down chart), which may require a single engine flyaway to be conducted below level of helipad. Lateral takeoff and approach to landing techniques used in Part A can be used whenever helipad is positioned with a vertical drop along one edge.

Part A for Elevated Helipads minimum (60 X 60 feet 18.3 X 18.3 meters)

·

*

*

Part B for Helipads (minimum 87 X 75 feet / 25.9 X 22.9 meters)

*

Part B allows for Category A operations from a ground level helipad. Takeoff and approach to landing techniques used in Part B can be utilized when there are obstacles that limit approach and departure path.

*

Part C allows for Category A operations using minimum distance possible conducted from a short runway.

Part C for Short Airfields (minimum 625 feet /190.5 meters).

Limitations, performance used.

procedures, flight path, and are specific to type of helipad

Part A allows for Category A operations on an elevated helipad (WAT charts = 90 feet / 27.9 meters

DEFINITIONS

BALKED LANDING:

Discontinuance of a landing approach and initiation of a climb out. Category A balked landing capability following an engine failure is assured at or prior to LDP

*

Land and stop safely on intended landing area after arrival at TDP; or,

*

Climb out from point of failure and attain single engine forward flight prior to arrival at LDP.

CATEGORY A TAKEOFF:

CATEGORY A PROFILE:

Operation of helicopter in such a manner that if one engine fails at any time after start of takeoff, helicopter can:

Category A operations are defined in three takeoff segments and one landing profile.

*

*

Return to, and safely stop on takeoff area before arrival at TDP; or,

Continue takeoff, climb out, and attain single engine forward flight after arrival at TDP.

*

*

CATEGORY A LANDING:

*

Operation of helicopter in such a manner that if one engine fails at any point during approach, helicopter can:

*

Segment 1 (Takeoff) is defined as profile from takeoff to VTOSSand positive rate of climb. Segment 2 (VToss Climb/gradient) is defined as profile from VTOSSand positive rate of climb to 200 feet above takeoff surface and Vy.

Segment 3 (Vy Climb/gradient) is defined as profile of Vy, then Vy climb to 1000 feet above takeoff surface. Landing profile is defined as flight path from LDP to a normal landing. i


FAA APPROVED

COMPLETEDLANDING:

TAKEOFF SAFETY SPEED (VToss):

Procedure and profile from landing decision point (LDP) to complete stop on designated surface with one engine inoperative.

Airspeed that will assure required climb performances with one engine inoperative during Segment2 of CategoryA procedures.

DROPDOWNHEIGHT:

TARGETTORQUE:

Heightloss below takeoff surface height.

Available 21/2minute OEI torque of minimum specification engine, assured by satisfactory completion of power assurancecheck.

COMPLETED TAKEOFF:

During CategoryA takeoff, engine failure after Takeoff Decision Point (TDP), continued takeoff is assured.

NOTE

All charts are based upon minimum specification enginepower.

FACTOREDWINDS:

Unfactoredwind speed decreasedby amount specified by regulatoryagency(i.e: DOT/ FAA requires headwind to be factored by 50%. A reported headwind of 26 knots factored by 50% will result on 13 knots to be used in performancecalculations).

UNFACTORED WINDS:

Wind speed and direction as reported by tower or weather office.

LANDINGDECISIONPOINT(LDP):

Last point in landing path where it is still possible for helicopter to reach VTOSS and initiate a go around (balked landing).

ABBREVIATIONS

LANDINGDISTANCE REQUIRED:

Horizontal distance from a point after LDP where helicopter is above landing surface to point where it is brought to a safe stop on designatedsurface with one engine operative. REJECTED TAKEOFF:

During CategoryA takeoff, enginefailure prior to Takeoff Decision Point (TDP), completed landing backto takeoff surface is assured. TAKEOFFDECISION POINT(TDP):

Last point in takeoff path where a rejected takeoff can beassured, andfirst point wherea completed takeoff can be assured.

FP

-

Flying Pilot

LDP NFP OEI TDP

-

Landing Decision Point

-

Non Flying Pilot

-

One Engine Inoperative

VROT

-

- Takeoff Decision Point

Velocity at which airspeed

indicator has perceivablemotion

VTOSS -Takeoff

Safety Speed

Vy

-Best Rateof Climb Speed

WAT

- Weight-altitude-temperature

T Section A RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4

FAA APPROVED

PART A - ELEVATED HELIPAD

1A-3. TYPES OF OPERATION

Category A operation is approved for day/ night VMC, non-icing conditions. Night takeoff and landing may be accomplished with adequatelighting.

Flight shall be planned to include alternate OEI landing site, if OEI,OGEhover capability is not available( SeeManufacturer'sData). Training of actual single engine elevatedpad landing or takeoff is prohibited if operation over continuous OEIrange is required. Minimum of 15 feet (4.6 meters) obstacle clearance must be maintained throughout initial takeoff path.

Helipad shall be positioned so that one edge is within 30 feet (9.1 meters)of a vertical drop for emergency flyaway in event of an engine failure. Minimumhelipaddimensionsare 60feet by 60 feet (18.3meters by 18.3meters).

1A-4. FLIGHT CREW

Minimum crew for Category A elevated helipad operation consists of two pilots. Either crew position may be utilized by pilot flying helicopter.

0

1A-5. CONFIGURATION

Helicopter may be equippedwith standardor high skids andemergencyfloats (If required). Approved radar altimeter indicator(s) in view of both pilots.

Dual controls and copilot instrument kits. Operational intercommunication

system

(ICS).

Operation and performancevalues are based on all doors closed and standardheateroff.

1A-6. WEIGHT AND CENTER OF GRAVITY

1A-6-C. TAKEOFF AND LANDING WEIGHT-ALTITUDETEMPERATURE LIMITATIONS

If helipad height is 90 feet (27.4 meters) or greater, refer to Weight-Altitude-Temperature Limits chart (Figure 1A-1). Chart assures 15 feet (4.6 meters) obstacleclearance. If helipad height is less than 90 feet (27.4 meters), refer to Drop down height chart (Section 4A) to determine allowable gross weight and obstacle clearance.

1A-6-D. WINDAZIMUTH/GROSS WEIGHTBENEFIT

Wind for takeoff or landing shall be within 45°° of helicopter headingon landing side and 90 on takeoff side (see Section2A). Maximum crosswind component not to exceed10 knots (see Figure 1A-2).

° Headwind benefit only allowed if wind +45 of

takeoff/landingdirection.


FAA APPROVED

1A-7. AIRSPEED

NOTE

Unless otherwise authorized by operating regulations, pilot Is not authorized to credit performance increaseresulting from more than 50% of unfactoredheadwindcomponentfor takeoff segmentor landing.

1A-7-A. CLIMB AIRSPEEDS

Derived maximum GW for Category A Operations is corrected by Weight-AltitudeTemperature Limits chart (Figure 1A-1) and Headwind and Crosswind Component chart (Figure 1A-2). EXAMPLE:

(Enginemodel PT6T-3D).

VTOSS

40 KIAS

Vy

70 KIAS

1A-10.

HEIGHT-VELOCITY

No height-velocity limitationsexist if Category A takeoffs and landings are conducted within flight profile and appropriate Weight-AltitudeTemperature(WAT)chart.

1A-13.

POWER PLANT

Hp - 2000ft.

Category A operations are prohibited if installed engines do not meet specification power.

OAT - 30°C.

NOTE

Refer to power assurance chart in PERFORMANCESection 4 of BHT-412-FM-4 or BHT-412-FMS-46.3 for PT6T-3Dengines.

Upon entering unfactored headwind portion of WATchart, follow guidelines down to factored wind or temperature limit line, whichever occurs first. Then proceed down to determinetotal gross weight allowed.

Refer to power assurance chart in PERFORMANCESection 4 of BHT-412-FMS53.4for PT6T-3DE engines.

Allowable zero wind gross weight -

10,300pounds. Actual

unfactored

wind

(± 30° off

helicopternose) - 20 knots. Crosswind component -

10 knots.

Unfactored headwind component

-

17.5knots.

Factored wind (50% x 17.5 knots) -

8.75knots.

Weight benefit from factored wind 450pounds. Allowable gross weight pounds.

10,750

Refer to power assurance chart in PERFORMANCESection 4 of BHT-412-FMSengines. 56.3 & 56.4for PT6T-3DF

1A-13-B. POWER TURBINE RPM

Performance and procedures are predicated on P&W PT6T-3D,-3DE,and -3DF engines being installed utilizing ITT OEI Limit of 940C. ENGINEPOWERTURBINERPM(Np)shall be 103%, during takeoff and approaches only, and airspeed not exceeding80 KIAS.


FAA APPROVED

WEIGHT-ALTITUDE-TEMPERATURE LIMITS ELEVATEDHELIPAD TAKEOFF AND LANDING

PT6T-3DENGINE 103%RPM(AEO) GENERATOR 150AMPS

36

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

HEATEROFF

52

54

ENTER

412FS62.3-1A1-1

Figure 1A-1. Weight-Altitude-Temperature Limits (Sheet 1 of 3)

3


BHT-412-FMS-62.3& 62.4


PT6T-3DEENGINE 103%RPM(AEO) GENERATOR150AMPS

HEATEROFF

GROSS WEIGHT-100 KG

36

Figure 1A-1

38

40

42

44

46

48

50

52

54

Weight-Altitude-Temperature Limits (Sheet 2 of 3)


WEIGHT-ALTITUDE-TEMPERATURE LIMITS ELEVATED HELIPAD TAKEOFFAND LANDING

PT6T-3DFENGINE 103%RPM(AEO) GENERATOR 150AMPS

36

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

HEATER OFF

52

54

-2000

8000

9000

10000

11000

12000

GROSS WEIGHT - LB

0

10

TEMP

412FS62.3-1 A1-3

Figure 1A-1


5


BHT-412-FMS-62.3& 62.4

40

40

35

35

30

30

25

25

REPORTEDWIND SPEED- KNOTS

HEADWINDCOMPONENT- KNOTS

A 20 15

CROSSWINDANGLE FROMTAKEOFFHEADING

1.

3. 4. 5. 6. 7.

8.

EXAMPLE: TAKEOFFHEADING.....................................................................170

DIFFERENCEBETWEENWIND DIRECTIONAND TAKEOFFHEADING........................... REPORTEDWIND SPEED ............................................................... ENTERCHARTAT REPORTEDWIND SPEED, POINTA. PROCEEDLEFTHORIZONTALLY,FOLLOWINGSHAPEOF CURVEDLINES, TO CROSSWINDANGLE FROMTAKEOFFHEADING POINTB. PROCEEDVERTICALLYUPWARDTO CROSSWIND COMPONENTLIMITSCALE AND READCROSSWINDCOMPONENT............................ (NOTE: CHECKSECTION1A CROSSWINDLIMITATIONS) PROCEED LEFTHORIZONTALLYTO HEADWINDCOMPONENT SCALE AND READ UNFACTOREDHEADWINDCOMPONENT.................................. (NOTE: REDUCEHEADWINDCOMPONENTBY APPROPRIATEAPPROVEDFACTOR)

Figure 1A-2. Headwind and Crosswind Component Chart

°

30 20 KNOTS

10 KNOTS

17.5 KNOTS

T N Section 2A I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-62.3& 62.4


2A-2. PREFLIGHTPLANNING All Category A takeoffs and landings should begin with thorough planning; elevated helipad landings should consider alternate landing options in event of an engine failure prior to LDP. Each segment, 1, 2, or 3 should be planned for by utilizing correct charts contained and identified within Section 4A. Each segment can be planned and actual climb gradient calculated to ensure necessary obstacle clearance is available.

Values derived from performance charts should be readily available to both pilots and reviewed, as necessary, just prior to takeoff.

2A-7. BEFORE TAKEOFF

2A-7-A. POWERASSURANCE CHECK

Power assurance check shall be accomplished satisfactorily prior to conducting Category A operations. (See LIMITATIONS - POWER PLANT)

2A-8. TAKEOFF/LANDING

2A-8-A. GENERALINSTRUCTIONS

Takeoff decision point (TDP) and landing decision point (LDP) are same. Visually determined, TDP and LDP point is 35 feet (10.7 meters) above and 35 feet (10.7 meters)

laterally from center line of helipad with proper sight picture viewed through flying pilot (FP) door window.

FP may occupy either crew seat to effect correct sight picture relative to takeoff and landing direction. FP should adjust seat so that TDP/LDP sight picture will be consistent. This position should provide FP with a consistent and proper field of view. Refer to Figures 2A-1 and 2A-2.

Flight crew coordination is paramount for elevated helipad takeoffs and landings. FP must concentrate fully on helicopter control while remaining focused on point of intended landing during takeoff or landing operations. Primary function of non-flying pilot (NFP) is to monitor engine and flight instruments, radar altimeter and any other parameters deemed necessary. NFP will also have hand on collective in preparation to assist FP in adjustment of power to 2 1/2 minute OEI power setting in event of an emergency flyaway.

2A-8-B. TAKEOFF

NOTE

Refer to Normal takeoff profile (elevated helipad) (Figure 2A-1) and Decision points (TDP or LDP) sight pictures (Figure 2A-2). Takeoff shall be initiated with helicopter positioned such that takeoff sight picture is directly out FP door window.

7

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4 BRIEFING:

sight picture at TDP.

Review any factors which may affect Segment1 (Takeoff).

FP - Control flight path by visual reference and focus on point of intendedlanding.

DetermineFP and NFPduties.

FP - Upon reaching TDP positioning, call FLYAWAY and applies collective to accelerate to Vy airspeed.

Determine if wind direction and velocity are as forecast. Verify maximum gross weight for CategoryA procedure.

Determine 2 1/2minute power (Target Torque).

Seats -

FAA APPROVED

Adjusted.

FP - After obtaining Vy and 200 feet (61 meters)above helipadadjust Np to 100%prior to accelerationpast 80 KIAS.

2A-8-C. LANDING

Collective- Full down (flat pitch). Altimeter -

NFP- Monitorengine parameters.Prior to Vy assist FPduring emergencies.

Set.

Radar altimeter - Set decision height indicator to 35 feet (10.7 meters) plus helipad elevation.

BRIEFING:

Review any factors which may affect landing. DetermineFP seat (i.e. left or right).

ALTNSTATICSOURCEswitch - PRI.

ReviewFP and NFPduties.

RPMINCR/DECR switch - 103%Np

Verify maximum gross weight for condition, if different than flight planned.

Other instruments ranges.

Normal operating

Area - Clear.

Hover - Approximately 4 feet (1.4 meters) skid height. Main rotor torque - Noted.

NFP - Monitor instruments, advise FP on reaching limits, keep left hand on collective, report radar altitude, inform FP in case of malfunctions,and emergencies.

Collective- Apply smoothly to obtain a slow lateral climb to a point 35 feet (10.7 meters) above and alongside helipad with proper

Determine 21/2 minute power (Target Torque).

Seats - Adjusted.

Flight controls - Adjust friction to desired level.

Throttles -

Full open.

RPM INCR/DECR switch airspeedbelow 80 KIAS. Other instruments ranges.

103% Np at

Normal operating


BHT-412-FMS-62.3 & 62.4

Altimeter - Set.

NFP - Monitor engine parameters, keep left hand on collective, report radar altitude, inform FP in case of limitations, malfunctions, and emergencies.

height - Set decision Radar altimeter indicator to 35 feet (10.7 meters) plus helipad elevation. ALTN STATIC SOURCE switch - PRI. Flyaway area - Clear.

FP - Upon reaching LDP position, call LAND and slowly descend laterally onto helipad while focusing on point of intended landing. NFP -

Monitor instruments, report altitudes.

FP - Conduct normal approach to hover point (LDP) 35 feet (10.7 meters) above and alongside helipad with proper sight picture of helipad through FP door window.

9


FAA APPROVED

V

35 FT

HELIPAD HEIGHT

TDP/LDP

ROTATED 90 FOR CLARITY

Figure 2A-1 Normal TakeoffProfile (elevated helipad)


BHT-412-FMS-62.3

/

\

LANDING TO RIGHT

RIGHT CREW WINDOW

& 62.4

/

LANDING TO LEFT

LEFT CREW WINDOW

412FS62.3-2A-2

Figure 2A-2 Decision Points (TDP or LDP) Sight Pictures

11

T N I Section 3A R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O & 62.4 BHT-412-FMS-62.3

FAA APPROVED


3A-3. ENGINE

3A-3-A. SINGLE ENGINE FAILURE (MALFUNCTIONS RESULTING IN LOSS OF POWER: I.E. ENGINE FAILURE, LOW SIDE GOVERNOR FAILURE)

Engine failure can occur before, at, or after TDP/LDP.Since TDPand LDP are defined as same point, an engine failure at TDP/LDPcan result in either a FLYAWAYor LANDING,as committedto by pilot. 3A-3-A-1. DURING TAKEOFF PRIOR TO TDP (REJECTEDTAKEOFF)

Refer to Figure 3A-1 Emergency Landing/ TakeoffProfile.

An engine failure prior to reaching TDP (35 feet (10.7 meters) above and alongside heliport with proper sight picture) will necessitatea landingback to heliport.

NFP inform FP at indication of an engine failure and monitor engine parameters and ROTORRPMand alert FP if any limitations are reached.

Landingis accomplishedby descending back toward helipad while maintaining ROTOR RPM. While maintaining a level attitude, increase collective as necessary to cushion landingand eliminatelateral drift using cyclic. Performnormal engineshutdown.

3A-3-A-2. DURING TAKEOFF AT OR AFTER TDP (COMPLETED TAKEOFF)

Refer to Figure 3A-1 Emergency Landing/ TakeoffProfile. In event of an engine failure at TDP and FP proceed as follows: commits to FLYAWAY, safe clearancefrom helipad FP shall maintain by utilizing a slight lateral cyclic slide and yawing helicopter 10/15° away from Helipad and accelerate(10 to 15degrees nose down) helicopter to VROTwhile descending below level of helipad. At VROT, FP will rotate, continue acceleration to VTOSS (40 KIAS) and begin climb. NFPwill assist FP in adjusting engine power to 2 1/2 minute OEI power (TargetTorque) setting. FP will maintain2 1/2minuteOEI power setting to VTOSS and 200feet (61meters) above helipad. Upon reaching 200 feet (61 meters) above helipad,accelerateto Vy (70 KIAS). At end of 2 1/2 minute OEI limit, decrease power to 30 minuteor MCPpower setting and 100%/ NP / NR. At safe altitude, apply OEI procedures (Section 3, SINGLE ENGINE FAILUREinflight BasicFlight Manual). 3A-3-A-3. DURING LANDING PRIOR TO OR AT LDP (BALKED LANDING) Refer to Figure 3A-1 Emergency Landing/ TakeoffProfile.

In event of an engine failure prior to LDP,FP shall maintain safe clearance from helipad and accelerate (10 to 15 degrees nose down) helicopter to VROT, rotate, continue


acceleration to VToss and begin climb. NFP will assist FP in adjusting engine power to 2 1/ 2 minute OEI power (Target Torque) setting. FP will maintain 2 1/2 minute OEI power setting while climbing to 200 feet (61 meters) above helipad and accelerating to Vy (70 KIAS). At end of 2 1/2 minute OEI limit, decrease power to 30 minute or MCP power setting and 100% / NP / NR. At safe altitude, apply OEI procedures (Section 3, SINGLE ENGINE FAILURE INFLIGHT in Basic Flight Manual) and proceed to alternate OEI landing site. 3A-3-A-4.

DURING LANDING AT OR AFTER LDP (COMPLETED LANDING)

Refer to Figure 3A-1 Emergency Landing/ Takeoff Profile.

An engine failure after reaching LDP (35 feet (10.7 meters) above and alongside heliport with proper sight picture) will necessitate a landing to heliport.

NFP inform FP at indication of an engine failure and monitor engine parameters and ROTORRPM and alert FP if any limitations are reached.

BHT-412-FMS-62.3& 62.4

Landing is accomplished by descending toward helipad while maintaining ROTOR RPM. While maintaining a level attitude, increase collective as necessary to cushion landing and eliminate lateral drift using cyclic. Perform normal engine shutdown. 3A-3-A-5.

WARNING/CAUTION LIGHTS (ANY MALFUNCTION THAT DOES NOT RESULT IN ENGINE POWER LOSS)

Control rotor RPM. Before TDP:

If Master Caution light illuminates, perform rejected takeoff and take appropriate corrective action (see Basic Flight Manual).

After TDP:

If Master Caution light illuminates, continue takeoff profile. After obtaining VY and 200 feet altitude, take corrective action (see Basic Flight Manual).

At or after LDP:

If Master Caution light illuminates, continue approach to landing and take appropriate action (see Basic Flight Manual).

13


FAA APPROVED

Vy (70 KIAS)

RATE OF CLIMB

TAKEOFF

412FS62.3-3A-

Figure 3A-1

Emergency Landing/Takeoff Profile

T N I Section4A R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4

FAA APPROVED

PART A -

ELEVATED HELIPAD

4A-1. INTRODUCTION

Performance data presented in this supplement are based on engine manufacturer'sminimum specification power with installation losses.

4A-2. POWERASSURANCE CHECK Refer to power assurance

check chart in

Basic Flight Manual Section 4 PERFORMANCE, or appropriate FMS for installed engines (see LIMITATIONS - POWER PLANT). Instructions for chart use can be found on chart. Power assurance check chart is based on a minimum specification engine with installation losses and proper rigging of enginecontrols.

If engine performance does not meet that shown in power assurancecheck, CategoryA operations will not be attempted and steps must be taken to ascertain causes of engine power loss.

4A-6. TAKEOFF DISTANCE/

PROFILE

Elevated helipad takeoff distance/profile begins at TDP /LDP (Figure 2A-1) and proceeds to a point 1000feet (305 meters) abovetakeoff surface.This distance and path, after engine failure, can be defined in three segments.

Segment 1 (Takeoff):

WAT Chart - maximum allowable gross weight (Figure 4A-1.).

Headwind/crosswindcomponent chart (Figures4A-2).

2 1/2minute OEI rating (Figure 4A-3).

Drop down height (Figure 4A-4). Takeoff distance/path

VTOSS and positive

from

TDP to

rate of climb

(Figure 4A-5).

Segment2 (VToss climb/gradient):

Distance/pathfrom VTOSSto 200 feet (61 meters) above takeoff surface (Figure 4A-6).

Distance for level acceleration from VTOSSto Vy (Figure 4A-7).

Segment3 (Vy climb/gradient):

Distance/pathfrom 200 feet(61 meters) to 1000 feet (305 meters) above takeoff

surface (Figure 4A-8).

EXAMPLES.ELEVATEDHELIPAD

What is takeoff distance/profile following conditions?

given

Engines - PT6T-3DF. Altitude -

Sea level.

Temperature - 23°C. Wind -

15 KTS + 35° to helicopter

nose.

HelipadHeight - 90 feet (27.4meters). 15-


FAA APPROVED

Step 1. Wind benefit:

meters) clearance.

°

15 knots - within wind limitation ± 45 of helicopter nose (Figure 4A-2). Headwind component -

12 knots.

NOTE

Unless otherwise authorized by operating regulations, pilot is not authorized to credit performance increase resulting from more than 50% of unfactored headwind component for takeoff segment or landing. Factored headwind component - 6 knots.

Cross wind component - 8.5 knots.

Step 2. Maximum allowable gross weight (Figure 4A-1): Zero wind weight-

11,250 pounds.

NOTE

Upon entering unfactored headwind portion of WAT chart, follow guide lines down to factored wind or temperature limit line, whichever occurs first. Then proceed down to determine total gross weight allowed. Headwind benefit - 250 pounds.

Total -

11,500 pounds.

Step 3. OEI (Target) torque:

2 '/2minute rating (Figure 4A-3) - 67%.

Step 4. Drop down height:

Height loss (Figure 4A-4) (use zero wind gross weight) - 78 feet (23.8 meters) height loss from TDP or 43 feet (13 meters) drop down below takeoff surface. Obstacle clearance 90 feet (27.4 meters) helipad height - 43 feet (13 meters) drop down, 47 feet (14.3

Step 5. Takeoff (Segment 1):

Distance from TDP to VTOSS (Figure 4A-5)- 500 feet (152.4 meters).

Step 6. VTOSS Climb (Segment 2):

Distance and climb gradient VTOSS to 200 feet (61 meters) above takeoff surface - (Figure 4A-6). 13% climb gradient.

43 feet (13 meters) (drop down) + 200 feet (61 meters) - .13 - 1,869 feet (570 meters). VTOSS to VY level acceleration distance (Figure 4A-7)- 1,270 feet (387 meters).

Step 7. Vy Climb (Segment 3):

Distance from 200 feet (61 meters) to 1,000 feet (305 meters) above takeoff surface area (Figure 4A-8). 8.5% climb gradient.

800 feet (244 meters) -. 085 - 9,412 feet (2868 meters).

ANSWER:

Total Takeoff Distance Step 5

-

500

+Step 6

-

3,139

+Step 7

-

9,412

13,051 Feet (3,978 Meters)

4A-7. CLIMB AND DESCENT

4A-7-B. SINGLE ENGINERATE OF CLIMB CHARTS These charts and their locations are provided for additional information.


FAA APPROVED

where a single engine landing (see Manual) can be Basic Flight accomplished.

Vy (70 KIAS):

See Basic Flight Manual or appropriate (see engines FMS for installed LIMITATIONS- POWER PLANT).

For planning purposes, use twin engine fuel flow charts (Expanded Performance, Manufacturer's Data) for single engine cruise fuel These charts are requirements. conservative when used for OEI fuel requirements, since specific range for OEI cruise is always greater than twin engine.

VTOSS(40 KIAS):

Figure 4A-9 is a function of gross weight, H, OAT, and is same for PT6T3D, DE, and DF engines at 21/2 minute power.

4-A-7-C. ALTERNATE OEI LANDING SITE

When planning a flight terminating in an elevated helipad landing, and a single engine failure occurs prior to reaching destination; two options are available: 1.

2.

If fuel consumed enroute to original destination allows sufficient gross weight reduction to OEI hover capability, a single engine landing may be accomplished using normal landing profile (OEI Hover Chart, Expanded Performance, Manufacturer's Data).

Sufficient fuel reserve should be planned for a suitable alternate

17


FAA APPROVED



36

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

HEATEROFF

52

54

Figure 4A-1. Weight-Altitude-Temperature Limits (Sheet 1 of 3)


BHT-412-FMS-62.3& 62.4

LIMITS WEIGHT-ALTITUDE-TEMPERATURE

PT6T-3DEENGINE 103%RPM(AEO) 150 AMPS GENERATOR

Figure 4A-1

ELEVATEDHELIPAD TAKEOFF AND LANDING

HEATEROFF


19


FAA APPROVED

WEIGHT-ALTITUDE-TEMPERATURE LIMITS ELEVATED HELIPAD TAKEOFF AND LANDING

PT6T-3DFENGINE 103%RPM(AEO) GENERATOR150AMPS

36

38

GROSSWEIGHT -100 KG 40 42 44 46 48 50

HEATEROFF

52

54

Figure 4A-1 Weight-Altitude-Temperature Limits (Sheet 3 of 3)


FAA APPROVED

Figure 4A-2.

Headwind and Crosswind Component Chart

21


FAA APPROVED

POWER AVAILABLE

HEATER OFF

PT6T-3D/DE/DF ENGINE 100% RPM (OEI) GENERATOR 150 AMPS

Figure 4A-3

22

Power Available 2 1/2 Minute Rating

T N I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

PT6T-3D/DE/DF 2.5 MIN POWER 100% RPM (OEI) GENERATOR150 AMPS

BHT-412-FMS-62.3 & 62.4

DROP DOWN HEIGHT ELEVATED HELIPAD

HEATEROFF

NOTE: Use Zero Wind Weight for determination of Drop Down Height

HEIGHT LOSS = 35 FT + DROP DOWN HEIGHT

NOTE: FOR DEN ALT LESS THAN 1000 FT USE 1000 FT LINE

-40

-20

0

20

40

OAT - DEG C

60 8.0

8.5

9.0

9.5

10.0

10.5

11.0

11.5

12.0 0

GROSSWEIGHT-1000 LB

20

40

60

80

100

DROP DOWN HEIGHT - FT

Figure 4A-4.

120 140 412FS62.3-4A-4

Drop Down Height Elevated Helipad

23

T N I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4

FAA APPROVED

SINGLE ENGINETAKEOFF DISTANCE ELEVATEDHELIPAD

PT6T-3D/DE/DF ENGINE 2.5 MIN POWER 100% RPM (OEI) GENERATOR 150 AMPS

HEATER OFF

0

-40

-20

0

20

TAKEOFF DISTANCE - FT *MINIMUM DISTANCE 500 FEET

ENGINE FAILURE AFTER TDP

TDP

TAKEOFF DISTANCE

VROT

Figure 4A-5

24

Single Engine Takeoff Distance Elevated Helipad

412FS62.3-4A-5


FAA APPROVED

TAKEOFF FLIGHT PATH FROM VTOSS TO 200 FEET ABOVE TAKEOFFSURFACE

PT6T-3D/DE/DF ENGINE 2.5 MIN POWER 100% RPM (OEI) GENERATOR

150 AMPS 6000 5000

4000

7000

0

Figure 4A-6. Takeoff Flight Path From VTOSSto 200 Feet (61 Meters) Above Takeoff Surface

25

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O I-

400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 HORIZONTAL DISTANCE - FEET

412FS62.3-4A-7-1 Figure

26

4A-7.

Acceleration

VTOSS To Vy (Sheet

1 of 3)


BHT-412-FMS-62.3 & 62.4

HORIZONTALDISTANCEFOR LEVELACCELERATION

PT6T-3DEENGINE 30 MINUTE POWER 100% RPM (OEI) GENERATOR150 AMPS

FROMVTOSSTO VY ZERO WIND

HEATEROFF

HORIZONTAL DISTANCE - METERS 150 200 250 300 350 400 450 500 550 600 650 700 750 800

Figure 4A-7.

Acceleration VTOSSTo Vy (Sheet 2 of 3)

27


BHT-412-FMS-62.3 & 62.4


ENGINE PT6T-3DF 30 MINUTEPOWER 100%RPM(OEI) 150AMPS GENERATOR

FROMVTOSSTO VY ZEROWIND

HEATEROFF

8000

412FS62.3-4A-7-3

Figure 4A-7.

Acceleration

VTOSSTo Vy (Sheet 3 of 3)


BHT-412-FMS-62.3& 62.4

TAKEOFF FLIGHT PATH200 TO 1000 FEET ABOVE TAKEOFFSURFACE

PT6T-3DENGINE

VY =

100%RPM(OEI) GENERATOR 150 AMPS

62.3-4A8-1

Figure 4A-8

Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 1 of 3)

29


BHT-412-FMS-62.3& 62.4

TAKEOFF FLIGHTPATH200 TO 1000 FEET ABOVE TAKEOFF SURFACE

PT6T-3DE ENGINE 30 MINUTE POWER 100% RPM (OEI) GENERATOR 150 AMPS

Figure 4A-8 Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 2 of 3)


FAA APPROVED

TAKEOFF FLIGHT PATH 200 TO 1000 FEET ABOVE TAKEOFF SURFACE

VY= 70 KIAS HEATEROFF

PT6T-3DFENGINE 30 MINUTEPOWER 100%RPM(OEI) GENERATOR150AMPS

1600

5000 4000

1000

3000

1000

400

5000

11000 5000

400

5400

4200 4200 4400 00

8000

3200

7000

0

4

8

12

16

20

24

28

CLIMB GRADIENT- %

32

36

40

412FS62.3-4A8-3

Figure 4A-8. Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 3 of 3)

31


FAA APPROVED


PT6T-3D/DE/DF ENGINE 2.5MINUTE 100%RPM(OEI) GENERATOR 150AMPS

-1

0

1

VTOSS= 40KIAS HEATER OFF

2

RATE OF CLIMB - M/SEC 3 4 5 6 7

8

9

10

11


412FS62.3-4A-9

Figure 4A-9.

32

VTOSS(40 KIAS) Climb

T N I Section B R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4

FAA APPROVED

PART B -

GROUND LEVEL HELIPAD

1B-3. TYPES OF OPERATION

Category A operation is approved for day/ night VMC, non-icing conditions. Night takeoff and landing may be accomplished with adequate lighting. Flight shall be planned to include alternate OEI landing site (if required). If OEI is experienced prior to LDP, landing may be accomplished at weights at least 500 Lbsless than WAT Chart values (Figure 1B-1). See Emergency Procedure (Alternate OEI Landing, Section 3B). Training of actual single engine ground level helipad landing or takeoff is prohibited if operation over continuous OEI range is required. Minimum of 15 feet (4.6 meters) obstacle clearance must be maintained throughout initial takeoff path.

Minimum helipad -87 X 75 feet (25.9 X 22.9 meters) with 25 feet (7.6 meter) clear area around helipad in takeoff / landing direction (see Section 2B).

1-B-4. FLIGHT CREW

Minimum crew for Category A ground level helipad operation consists of one pilot operating helicopter from right seat. If dual pilot, either crew position may be utilized by flying pilot.

1B-5. CONFIGURATION

Helicopter may be equipped with standard or high skids and emergency floats (if required).

1B-6. WEIGHT AND CENTER OF GRAVITY 1B-6-C. TAKEOFF AND LANDING WEIGHT-ALTITUDETEMPERATURE LIMITATIONS Refer to Weight-Altitude-Temperature Limits chart (Figure 1B-1). Use Zero Wind Weight for determination Drop Down Height (Refer to Section 4B).

of

NOTE

WAT charts insure minimum obstacle clearance of 15 feet. If obstacle clearance other than 15 feet are necessary, then use Drop Down Height chart (Section 4B) to determine allowable gross weight.

1B-6-D. WIND AZIMUTH/GROSS WEIGHT BENEFIT

Wind for takeoff or landing shall be within +90 ° of takeoff/landing direction. Maximum crosswind component not to exceed 10 knots (see Figure 1B-2).

° Headwind benefit only allowed if wind ±45 of takeoff/landing direction.

Approved radar altimeter indicator in view of pilot. Operation and performance values are based on all doors closed and standard heater off. 33


BHT-412-FMS-62.3 & 62.4

1B-7. AIRSPEED

NOTE

Unless otherwise authorized by operating regulations, pilot is not performance authorized to credit increase resulting from more than 50% of actual wind for takeoff segment or landing.

Derived maximum GW for Category A Operations is attained by using Headwind and Crosswind Component chart and WAT chart (Figures 1B-2 and 1B-1).

1B-7-A. CLIMB AIRSPEEDS VTOSS

40 KIAS

Vy

70 KIAS

1B-10.

HEIGHT-VELOCITY

No height-velocity limitations exist if Category A takeoffs and landings are conducted within flight profile and appropriate Weight-Altitudetemperature (WAT) chart.

EXAMPLE:

(Engine model PT6T-3DE). HP - 2000 feet.

1B-13.

OAT - 25°C.

NOTE

Upon entering unfactored headwind portion of WAT chart, follow guide lines down to factored wind or temperature limit line, whichever occurs first. Then proceed down to determine total gross weight allowed. Allowable zero wind gross weight 10,000 pounds. ° Actual unfactored wind (± 10 off helicopter nose) - 30 knots.

Crosswind component - 5 knots.

Unfactored headwind component - 29 knots. Factored wind (50% x 29 knots) knots.

Category A operations are prohibited if installed engines do not meet specification power.

Refer to power assurance chart in PERFORMANCE Section 4 of BHT-412-FM-4 or BHT 412-FMS-46.3 for PT6T-3D engines.

chart in assurance Refer to power PERFORMANCE Section 4 of BHT-412-FMS53.4 for PT6T-3DE engines. Refer to power assurance chart in PERFORMANCE Section 4 of BHT-412-FMS56.3 & 56.4 for PT6T-3DF engines.

1B-13-B. POWERTURBINERPM

14.5

Weight benefit from factored wind 1,100 pounds. Allowable gross weight - 11,100 pounds.

POWER PLANT

Performance and procedures are predicated on P&W PT6T-3D,-3DE, and -3DF engines being Installed utilizing ITT OEI Limit of 940°C.

ENGINE POWER TURBINE RPM (Np) shall be 103%, during takeoff and approaches only, and airspeed not exceeding 80 KIAS.


FAA APPROVED

WEIGHT-ALTITUDE-TEMPERATURE LIMITS GROUND LEVEL HELIPAD TAKEOFF AND LANDING

PT6T-3DENGINE 103%RPM(AEO) GENERATOR150AMPS

36

50 00

-

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

HEATEROFF

52

54

-

Figure 1B-1. Weight-Altitude-Temperature Limits (Sheet 1 of 3)

35


FAA APPROVED

WEIGHT-ALTITUDE-TEMPERATURE LIMITS GROUNDLEVEL HELIPAD TAKEOFFAND LANDING

PT6T-3DEENGINE 103%RPM(AEO) GENERATOR 150 AMPS

36

Figure 1B-1.

36

38

GROSS WEIGHT-100 KG 40 42 44 46 48 50

Weight-Altitude-Temperature

HEATEROFF

52

54

Limits (Sheet 2 of 3)


BHT-412-FMS-62.3& 62.4

LIMITS WEIGHT-ALTITUDE-TEMPERATURE GROUND LEVEL HELIPAD

PT6T-3DF ENGINE 103% RPM (AEO) GENERATOR 150 AMPS

TAKEOFF AND LANDING


HEATER OFF


FAA APPROVED


CROSSWINDANGLE FROMTAKEOFFHEADING FROM TAKEOFF HEADING

1. 2. 3. 4. 5. 6. 7. 8.

EXAMPLE: TAKEOFFHEADING..................................................................... REPORTEDWIND DIRECTION............................................................ DIFFERENCEBETWEENWIND DIRECTIONAND TAKEOFFHEADING........................... REPORTEDWIND SPEED ............................................................... ENTERCHART AT REPORTEDWIND SPEED,POINTA. PROCEEDLEFT HORIZONTALLY, FOLLOWINGSHAPEOF CURVEDLINES, TO CROSSWINDANGLE FROMTAKEOFFHEADINGPOINTB. PROCEEDVERTICALLYUPWARDTO CROSSWIND COMPONENTLIMITSCALE AND READ CROSSWINDCOMPONENT............................. (NOTE: CHECKSECTION1A CROSSWINDLIMITATIONS) PROCEEDLEFT HORIZONTALLY TO HEADWINDCOMPONENT SCALEAND READ UNFACTOREDHEADWINDCOMPONENT................................... (NOTE: REDUCEHEADWINDCOMPONENTBY APPROPRIATEAPPROVEDFACTOR)

°

170 ° 180 10° 30 KNOTS

5 KNOTS

29 KNOTS

412FS62.3-1B-2

Figure 1B-2. Headwind and Crosswind Component Chart

38

T N I * Section2B R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

*

BHT-412-FMS-62.3 & 62.4

PART B-

GROUND LEVEL HELIPAD

2B-2. PREFLIGHT PLANNING

2B-8. TAKEOFF/LANDING

All Category A takeoffs and landings should begin with thorough planning; ground level helipad landings should consider alternate landing options in event of an engine failure prior to LDP. Each segment, 1, 2, or 3 should be planned for by utilizing correct charts contained and identified within Section 4A.

2B-8-A. GENERAL INSTRUCTIONS

Each segment can be planned and actual climb gradient calculated to ensure necessary obstacle clearance is available.

Values derived from performance charts should be readily available to flying pilot (FP) and reviewed, as necessary, just prior to takeoff.

2B-7. BEFORE TAKEOFF

2B-7-A. POWERASSURANCE CHECK

Power assurance check shall be accomplished satisfactorily prior to conducting Category A operations. (See LIMITATIONS - POWER PLANT, Basic Flight Manual or appropriate Supplement).

Flying pilot (FP) shall adjust seat position to affect correct sight picture for takeoff/landing. Takeoff decision point (TDP) and landing decision point (LDP) are same for any ambient condition.

Pilot shall determine maximum helicopter weight allowed, height of TDP/LDP, and single engine (target) torque available (see Section 4B).

TDP/LDP sight picture should provide pilot with a consistent and proper field of view. Refer to Figures 2B-1 through 2B-3.

Sight picture is paramount for takeoffs and landings. Pilot must maintain helicopter control while remaining focused on point of intended landing during takeoff or landing operations.

2B-8-B. TAKEOFF PLANNING TAKEOFF:

CONSIDERATIONS

FOR

Review any factors which may affect Segment 1 (Takeoff).

39


BHT-412-FMS-62.3& 62.4

Verify wind direction and velocity.

Verify correct gross weight for takeoff. Determine 2% minute power (Target Torque). Determine height of TDP.

Refer to normal takeoff profile (ground level helipad) (Figure 2B-1) and Decision points (TDP or LDP) sight pictures (Figure 2B-3). Takeoff shall be initiated with helicopter positioned such that takeoff sight picture is (FP) pilot's out flying directly windshield.

Collective - Full down (flat pitch). Altimeter - Set.

Radar altimeter - Set decision height indicator to TDP height.

RPM INCR/DECR switch -

Other instruments ranges. Area -

PRI.

103% Np

Normal operating

Clear.

Hover - Approximately 4 feet (1.4 meters) skid height. Mast torque -

With approximately 10° nose down and level flight, accelerate to Vy airspeed. After obtaining Vy and 200 feet (61 meters) above helipad adjust Np to 100% prior to acceleration past 80 KIAS.

2B-8-C. LANDING PLANNING LANDING:

CONSIDERATIONS

FOR

Review any factors which may affect landing. Verify wind direction and velocity.

Verify correct gross weight for landing. minute power (Target Determine 2 1/2 Torque). Determine height of LDP.

Seats - Adjusted.

Noted.

Collective - From a hover, gradually apply 10% additional mast torque to obtain slow vertical climb.

40

Backup climb and sight picture (Figure 2B-3) are maintained until reaching TDP (Figure 2B1). Upon reaching TDP,apply positive forward cyclic to transition to forward flight.

NOTE

ALTN STATIC SOURCE switch -

At twenty feet, transition to backup climb is begun by adding 5% additional mast torque to maintain proper sight picture with a reasonable rate of climb.

Flight controls - Adjust friction to desired level. Throttles -

Full open.


FAA APPROVED

RPM INCR/DECR switch airspeed below 80 KIAS.

-

103% Np at

Other instruments - Normal operating ranges. Altimeter -

Set.

Radar altimeter - Set decision height indicator to LDP height. ALTN STATIC SOURCE switch - PRI. Flyaway area - Clear.

NOTE

Intended landing helipad is approached so as to arrive with an acceptable wind azimuth at predetermined LDP height above helipad.

Conduct normal approach to landing decision point (LDP) above helipad with proper sight picture. Upon reaching LDP, a smooth transition to a moderate rate of descent is begun, while maintaining correct sight picture, to arrive over center of helipad with hover power.


Figure 2B-1. Normal Takeoff/Landing Profile

42

FAA APPROVED


FAA APPROVED

TAKEOFF DIRECTION

24

39

V

412FS62.3-2B-2

Figure 2B-2. Helipad Markings and Dimensions

43


FAA APPROVED

PILOT (RIGHT SEAT)

CO-PILOT (LEFT SEAT)

Figure 2B-3. TDP/LDP Sight Picture

44

412FS62.3-2B-3

T Section 3B RIN P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O & 62.4 BHT-412-FMS-62.3

FAA APPROVED

PART B - GROUND LEVEL HELIPAD

3B-3. ENGINE

3-B3-A. SINGLE ENGINE FAILURE (MALFUNCTIONS RESULTING IN LOSS OF POWER: I.E. ENGINE FAILURE, LOW SIDE GOVERNOR FAILURE)

Engine failure can occur before, at, or after TDP/LDP.Since TDPand LDP are defined as same point, an enginefailure at TDP/LDPcan result in either a FLYAWAYor LANDING,as committedto by pilot. 3B-3-A-1. DURINGTAKEOFF PRIOR TO TDP (REJECTEDTAKEOFF)

Referto Figure 3B-1EmergencyLanding/ TakeoffProfile.

An enginefailure prior to reachingTDPabove helipadwith proper sight picture will necessitatelanding back at helipad. Whenenginefails, pilot will set 2 1/2 min OEI power (targettorque) and allow helicopter to settle back toward helipad. A slight nose low attitude (forwardcyclic) will put helicopteron properflight path.

Touchdown on helipadshould be slightly nose high to minimize slide, add collective as requiredto cushion landing and apply cyclic inputs to eliminatelateral drift. Perform normal engineshutdown.

3B-3-A-2. DURING TAKEOFF AT OR AFTER TDP (COMPLETEDTAKEOFF)

Referto Figure 3B-1 EmergencyLanding/ TakeoffProfile.

In eventof an enginefailure at TDPand pilot proceedas follows: commitsto FLYAWAY,

Pilot shall acceleratehelicopter (10degrees while adjusting to 2 1/2 nose down) to VROT minuteOEI engine power (targettorque).At pilot will rotate,continue accelerationto VROT, (40 KIAS)and begin climb. VTOSS

Pilot will maintain2 1/2 minute OEI power and 200 feet (targettorque) setting to VTOSS (61meters)above helipad.Upon reaching200 feet (61meters) abovehelipad, accelerateto Vy (70 KIAS). At end of 2 1/2 minute OEI limit, decrease powerto 30 minuteor MCPpower setting and 100%/ Np / NR.At safe altitude, apply OEI procedures(Section3, SINGLEENGINE FAILURE,Basic Flight Manual). 3B-3-A-3. DURING LANDING PRIOR TO OR AT LDP (BALKED LANDING)

Referto Figure 3B-1 EmergencyLanding/ TakeoffProfile.

In eventof enginefailure prior to LDP,Pilot will acceleratehelicopter(10 degreesnose and begin climb. Pilot will down)to VTOSS maintain2 1/2 minuteOEI power (target

45


torque) setting while climbing to 200 feet (61 meters) above helipad and accelerating to Vy (70 KIAS).

At end of 2 1/2 minute OEI limit, decrease power to 30 minute or MCP power setting and 100% / Np / NR.At safe altitude, apply OEI procedures (Section 3, SINGLE ENGINE FAILURE, Basic Flight Manual) and proceed to alternate OEI landing site. 3B-3-A-4.

DURING LANDING AT OR AFTER LDP (COMPLETED LANDING)

Refer to Figure 3B-1 Emergency Landing/ Takeoff Profile.

An engine failure after reaching LDP above heliport with proper sight picture will necessitate landing at heliport.

When engine fails, pilot will set 2 1/2 min OEI power (target torque) and allow helicopter to settle toward helipad. A slight nose low attitude (forward cyclic) will put helicopter on proper flight path.

Touch down on helipad should be slightly nose high to minimize slide, add collective as required to cushion landing and apply cyclic inputs to eliminate lateral drift. Perform normal engine shutdown. 3B-3-A-5.

ALTERNATE OEI LANDING PRIOR TO LDP (COMPLETED LANDING)

In event of engine failure prior to LDP, a completed landing may be accomplished in one of two ways. 1.

Pilot may plan on returning to a runway for normal single engine landing (Section 3, SINGLE ENGINE

FAA APPROVED

2.

FAILURE, Basic Flight Manual). When helicopter gross weight is at least 500 pounds less than WAT chart values, landing may be performed at ground level helipad by adding 20 feet to LDP height and applying 21/2minute OEI power (Target Torque) as helicopter approaches proper glide path and sight picture. Approach to glide path will be more rounded as helicopter settles down desired slope with final touchdown as in previous procedures (3B-3-A-4).

3B-3-A-6.

WARNING/CAUTION LIGHTS (ANY MALFUNCTION THAT DOES NOT RESULT IN ENGINE POWER LOSS)

Control rotor RPM. Before TDP:


After TDP:

If Master Caution light illuminates, continue takeoff profile. After obtaining Vy and 200 feet altitude, take corrective action (see Basic Flight Manual).

At or after LDP:


I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O EMERGENCYLANDING/TAKEOFFPROFILE GroundLevel Helipad

n

NORMAL TAKEOFF PROFILE

.................

OPTIONAL OEI LANDING ADD 20 FT TO LDP

--

-

--

OEI TAKEOFF PROFILE

--

OEI LANDINGPROFILE

200 FT

35 FT

.

T N Section4B I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-62.3& 62.4

PART B - GROUND LEVEL HELIPAD

4B-1. INTRODUCTION


Performance data presented in this supplement are based on engine manufacturer's minimum specification power with installation losses.

WAT Chart - Maximum allowable gross weight (Figure 4B-1).

4B-2. POWERASSURANCE CHECK

2

Refer to power assurance check chart in Basic Flight Manual Section 4 PERFORMANCE,or appropriate FMS for installed engines (see LIMITATIONS - POWER PLANT). Instructions for chart use can be found on chart.

Power assurance check chart is based on a minimum specification engine with installation losses and proper rigging of engine controls. If engine performance does not meet that shown in power assurance check, Category A operations will not be attempted and steps must be taken to ascertain causes of engine power loss.

4B-6. TAKEOFF DISTANCE/ PROFILE

For ground level helipad, takeoff distance/ profile is horizontal distance along takeoff path, beginning with helicopter on pad and progressing to a point 1000 feet (305 meters) above takeoff surface. This distance and path, after engine failure, can be defined in three segments.

Headwind/Crosswind component (Figure 4B-2). 1/2 minute

power rating (Figure 4B-3).

Height of TDP/LDP (Figure 4B-4).

Height loss from TDP (Figure 4B-5).

Single engine takeoff distance 35 feet (Figure 4B-6).

Segment 2 (VToss climb/gradient):

Takeoff flight path from VTOSSto 200 feet (61 meters) (Figure 4B-7). Acceleration distance from VTOSSto Vy (Figure 4B-8).

Segment 3 (Vy climb/gradient):

Takeoff flight path from 200 feet (61 meters) to 1000 feet (305 meters) (Figure 4B-9).

EXAMPLES, GROUND LEVEL HELIPAD:

What is takeoff distance/path, given following conditions? Engines -

PT6T-3DE.

Altitude - 500 feet (914.4 meters) Hp Temperature -

40°C.

Wind - 30 knots on helicopter nose.


FAA APPROVED

Step 1. Wind benefit:

NOTE

30 knots on helicopter nose (Figure 4B-2). NOTE

by authorized Unless otherwise operating regulations, pilot is not authorized to credit performance increase resulting from more than 50% of unfactored headwind component for takeoff segment or landing. Factored headwind component knots.

Appropriate WAT chart insures 15 foot minimum obstacle clearance for this example (see takeoff profile on Figure 4B-6).

OEI takeoff distance to VTOSSand 35 feet (10.7 meters) (Figure 4B-6) -740 feet (250 meters).

Step 5.

Distance from VTOSSto 200 feet (61 meters) above takeoff surface (Figure 4B-7).

15


Step 2.

Maximum allowable (Figure 4B-1):

16% Climb gradient.

gross weight

165 feet (56.3 meters) -. 16 (315 meters).

Step 6.

Upon entering unfactored headwind portion of WAT chart, follow guide lines down to factored wind or temperature limit line, whichever occurs first. Then proceed down to determine total gross weight allowed. Headwind benefit (see note) - 850 pounds. Total -

10,750 pounds.

Step 3.

OEI (Target) torque (Figure 4B-3) 61.5%:

Step 4.

Takeoff (Segment 1):

Height of TDP (Figure 4B-4) (33.5 meters).

104 feet

1032 feet

VTOSSto Vy level acceleration (Figure 4B-8) - 2650 feet (808 meters).

Zero wind weight - 9,900 pounds. NOTE

VTOSSClimb (Segment 2):

Vy Climb (Segment 3) (Figure 4B-9):

4.5% climb gradient.

800 feet (244 meters) -. 045 feet (5418 meters).

17,778


ANSWER:

Total Takeoff Distance Step 4 -

740

+Step 5 -

3,682

+Step 6 -

17,778

22,200 Feet (6,767 Meters)

Height loss from TDP (Figure 4B-5) Not computed (see example, Section 1B).

49


4B-7. CLIMB AND DESCENT

4B-7-A. SINGLE ENGINERATE OF CLIMB CHARTS

FAA APPROVED

engine failure occurs prior to reaching destination; two options are available: 1.

These charts and their location are provided for additional information.

See Basic Flight Manual or appropriate FMS for installed engines (see LIMITATIONS - POWER PLANT).

VTOSS(40 KIAS):

This chart is a function of gross weight, Hp OAT, and is same for PT6T3D,DE,DF engines at 2 1/2 minute power.

reserve should be suitable alternate engine landing (see Manual) can be

For planning purposes, use twin engine fuel flow charts (Expanded Performance, Manufacturer's Data) for single engine cruise fuel requirements. These charts are conservative when used for OEI fuel requirements, since specific range for OEI cruise is always greater than twin engine.

Vy (70 KIAS):

Rate of climb chart (VToss - 40 KIAS) (Figure 4B-10).

Sufficient fuel planned for a where a single Basic Flight accomplished.

2.

After engine failure prior to reaching LDP, if fuel consumption allows gross weight reduction of 500 pounds less than WAT chart value (figure 1B-1), then alternate OEI landings can be accomplished (Alternate OEI Landing, Section 3B).

4B-9. LANDING DISTANCE

4B-7-C. ALTERNATEOEI LANDING When planning a flight terminating in an ground level helipad landing, and a single

Landing distance over a 50 feet obstacle for all allowable gross weights is 200 feet.


FAA APPROVED



36

8000

38

GROSS WEIGHT-100 KG 40 42 44 46 48 50

9000

GROSS

4B-1. Figure 413-1. Figure

10000

WEIGHT

11000

HEATEROFF

52

54

12000

-LB

Weight-Altitude-Temperature Weight-Altitude-Temperature

of 3) 3) (Sheet 11 of Limits (Sheet Limits

51


FAA APPROVED


PT6T-3DEENGINE 103%RPM(AEO) GENERATOR150AMPS

36

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

HEATEROFF

52

54



FAA APPROVED

LIMITS WEIGHT-ALTITUDE-TEMPERATURE GROUNDLEVEL HELIPAD TAKEOFFAND LANDING

PT6T-3DFENGINE 103%RPM(AEO) GENERATOR 150 AMPS

36

8000

38

GROSSWEIGHT-100 KG 40 42 44 46 48 50

9000

10000

11000

HEATEROFF

52

54

12000



Figure 4B-2. Headwind and Crosswind Component Chart

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O PT6T-3D/DE/DF ENGINE

412FS62.3-4B-3

Figure 4B-3. Power Available 2 1/2Minute Rating

55


FAA APPROVED

HEIGHT OF TDP/LDP FOR GROUND LEVEL HELIPAD

103% RPM (AEO) GENERATOR 150 AMPS

HEATER OFF

6000 FT. H

-40

-20

0

20

OAT - °C

D

40

60 20

40

60

80

100

120

140

HEIGHT OF TDIP/LDP - FT * MINIMUM HEIGHT 80 FEET

CHART REPRESENTS MINIMUM 15 FEET OF OBSTACLE CLEARANCE. IF ADDITIONALCLEARANCEIS REQUIRED,TDP/ LDPSHOULDBE ADJUSTED UPWARD.

HEIGHTLOSSCHART(FIG 4B-3)CAN BE USEDTO DETERMINE ACTUALOBSTACLECLEARANCEFORWEIGHTSLESS THAN MAXIMUMALLOWABLE (SEEWAT- FIG4B-1).

412FS62.3-4B-4

Figure4B-4. Height of TDP/LDPFor GroundLevel

56

T N I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O FAAAPPROVED

BHT-412-FMS-62.3& 62.4

HEIGHT LOSS FROM TDP GROUND LEVEL HELIPAD

PT6T-3D/DE/DF 2.5 MIN POWER 100%RPM (OEI) GENERATOR150 AMPS

HEATER OFF

NOTE:UseZero Wind Weightfor determinationof Drop Down Height

TE FORDENALT LESSTHAN 1000FT USE1000FT LINE

m

II

-40

-20

20 40 0 OAT - DEG C

60 8.0

8.5

9.0 9.5 10.0 10.5 11.0 11.5 GROSS WEIGHT - 1000 LB

12.0 0

60 80 100 120 140 20 40 HEIGHT LOSS FROM TDP - FT 412FS62.3-4B-5

Figure 4B-5. Height Loss FromTDP

57

T N I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4

FAA APPROVED

SINGLE ENGINE TAKEOFF DISTANCETO 35 FT ABOVE TAKEOFF SURFACE GROUND LEVEL HELIPAD

PT6T-3D/DE/DF

HEATER OFF

ENGINE

2.5 MIN POWER 100% RPM (OEI) GENERATOR 150 AMPS

-40

-20

0

20

1000

TAKEOFF DISTANCE - FT ENGINEFAILURE AFTER TDP 500 FEET

TDP

VTOSS POSITIVE RATE OF CLIMB

412FS62.3-4B-6

Figure 4B-6. Single Engine Takeoff Distance To 35 Feet (10.7 Meters) Above Takeoff Surface

58


FAA APPROVED

TAKEOFF FLIGHT PATHFROM VTOSS TO 200 FEET ABOVE TAKEOFFSURFACE

PT6T-3D/DE/DF ENGINE 2.5 MIN POWER 100% RPM (OEI) GENERATOR 150 AMPS 6000

5000

7000

5

Figure 4B-7. Takeoff Flight Path From VTOSSTo 200 Feet (61 Meters) Above Takeoff Surface

59


FAA APPROVED

HORIZONTALDISTANCEFORLEVELACCELERATION

PT6T-3D ENGINE 30 MINUTE POWER 100% RPM (OEI) GENERATOR150 AMPS


HEATER OFF

HORIZONTAL DISTANCE - METERS

400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 HORIZONTAL DISTANCE - FEET

412FS62.3-4B-8-1

60

T N I 0 R P N O R I O T F C U 0T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4

FAA APPROVED

HORIZONTALDISTANCEFOR LEVELACCELERATION FROMVTOSSTO VY ZERO WIND

PT6T-3DE ENGINE 30 MINUTE POWER 100% RPM (OEI) GENERATOR 150 AMPS

HEATER OFF


150

400

200 250

600

800

300

350

400 450

1000 1200

500

550 600

650

700

1400 1600 1800 2000 2200

750

800

2400 2600

2800

HORIZONTAL DISTANCE - FEET

412FS62.3-4B-8-2

Figure 4B-8.

To Vy (Sheet 2 of 3) Acceleration VTOSS

61


FAA APPROVED

HORIZONTALDISTANCEFORLEVELACCELERATION FROMVTOSSTO VY ZERO WIND

PT6T-3DF ENGINE 30 MINUTE POWER 100% RPM (OEI) GENERATOR 150 AMPS

HEATER OFF


300

350

400

450

500

550

600

650

700

750

80

8000

LL

6000 4000 2000

0

400

600

800

1000

1200 1400 1600 1800 2000 2200 HORIZONTAL DISTANCE - FEET

2400

2600

2800

412FS62.3-4B-8-3

Figure 4B-8.

AccelerationVTOSSTo Vy (Sheet 3 of 3)


BHT-412-FMS-62.3 & 62.4

TAKEOFFFLIGHTPATH200 TO 1000 FEET ABOVE TAKEOFF SURFACE

PT6T-3DENGINE MAXCONTINUOUS POWER 100%RPM(OEI) GENERATOR 150 AMPS

VY = 70 KIAS

HEATEROFF

- 1800 -

1600

- 1400

CLIMBGRADIENT-%

*

412FS62.3-4B-9-1

Figure 4B-9. Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 1 of 3)

63


BHT-412-FMS-62.3& 62.4


PT6T-3DEENGINE 30 MINUTEPOWER 100%RPM(OEI) 6000

5000

Figure 4B-9, Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 2 of 3)


FAA APPROVED


PT6T-3DFENGINE 30 MINUTEPOWER 100%RPM(OEI) 150 AMPS GENERATOR

0

VY = 70 KIAS HEATEROFF

4

8

12

16

20

24

28

32

36

40

CLIMB GRADIENT -%

Figure 4B-9. Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 3 of 3)

65


-4

-3

-2

412FS62.3-4B-10 Figure 4B-10.

66

Single Engine

Rate of Climb (VTosS - 40 KIAS)

T N I Section 1C R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4

FAA APPROVED

PART C - RUNWAY

1C-3. TYPES OF OPERATION

on all doors closed and standardheater off.

Category A operation is approved for day/ night VMC,non-icing conditions.

1C-6. WEIGHT AND CENTER OF GRAVITY

Training of actual single engine runway landing or takeoff is prohibited if operation over continuous OEI rangeis required.

Minimum runway length is 625 feet (190.5 meters)(SeeFigure 1C-1). Night takeoffs and landings may be accomplished with adequate lighting, which should consist of both edge and threshold lights/markings.

1C-4. FLIGHT CREW

Minimum crew for Category A runway operation consists of one pilot operating helicopter from right seat. Landing and takeoff may be accomplished from left seat with dual pilots.

1C-5. CONFIGURATION

1C-6-C. TAKEOFFAND LANDING WEIGHT-ALTITUDETEMPERATURE LIMITATIONS

Referto Weight-Altitude-TemperatureLimits chart (Figure 1C-1).

1C-6-D. WIND AZIMUTH

No increase in total gross weight is allowed for wind benefit. Wind for takeoff or landing shall be within ° + 90 of takeoff/landing direction.

Maximum crosswind component not to exceed10 knots (see Figure 1C-2). EXAMPLE:

Helicopter may be equippedwith standard or high skids and emergencyfloats (if required). Approved radar altimeter indicator in view of flying pilot.

Operation and performancevalues are based

Actual wind - 20knots.

Wind direction - ± 30° off helicopter

nose.

Crosswind component- 10 knots.

67


FAA APPROVED

1C-7. AIRSPEED

Refer to power assurance chart in PERFORMANCE Section 4 of BHT-412-FMS53.4 for PT6T-3DEengines.

1C-7-A. CLIMB AIRSPEEDS VTOSS

40 KIAS

Vy

70 KIAS

1C-10. HEIGHT-VELOCITY

No height-velocity limitations exist if Category A takeoffs and landings are conducted within flight profile and appropriate Weight-AltitudeTemperature (WAT) chart.

1C-13. POWER PLANT

Category A operations are prohibited if power assurance check is unsatisfactory.

Refer to power assurance chart in PERFORMANCE Section 4 of BHT-412-FM-4 or BHT-412-FMS-46.3for PT6T-3D engines.

Refer to power assurance chart in PERFORMANCE Section 4 of BHT-412-FMS56.3 & 56.4 for PT6T-3DF engines.

1C-13-A. GAS PRODUCER RPM

Performance and procedures are predicated on P&W PT6T-3D,-3DE, and -3DF engines being installed utilizing ITT OEI Limit of 940°C.

1C-13-B. POWER TURBINE RPM

ENGINE POWER TURBINE RPM (Np) shall be 103%, during takeoff and approaches only, and airspeed not exceeding 80 KIAS.


WEIGHT-ALTITUDE-TEMPERATURE LIMITS RUNWAY TAKEOFF AND LANDING

PT6T-3DENGINE

HEATEROFF

103% RPM (AEO)

GENERATOR 150 AMPS

36

Figure 1C-1

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

Weight-Altitude-Temperature

52

54

Limits (Sheet 1 of 3)

69


FAA APPROVED

WEIGHT-ALTITUDE-TEMPERATURE LIMITS RUNWAY TAKEOFF AND LANDING

PT6T-3DEENGINE 103%RPM(AEO) GENERATOR 150 AMPS

HEATEROFF

36

38

GROSSWEIGHT-100 KG 40 42 44 46 48 50

52

54

5000

4000

412FS62.3-1C-1-2

Figure 1C-1 Weight-Altitude-Temperature Limits (Sheet 2 of 3) 70


LIMITS WEIGHT-ALTITUDE-TEMPERATURE RUNWAY TAKEOFF AND LANDING

HEATER OFF


36

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

52

54



FAA APPROVED

25


REPORTEDWIND SPEED- KNOTS

A 20

CROSSWINDANGLE FROMTAKEOFFHEADING

1. 2. 3. 4. 5. 6. 7.

8.

EXAMPLE: TAKEOFFHEADING..................................................................... REPORTEDWIND DIRECTION............................ ............................ DIFFERENCEBETWEENWIND DIRECTIONAND TAKEOFFHEADING................. REPORTEDWIND SPEED ............................................................... ENTERCHARTAT REPORTEDWINDSPEED, POINTA. PROCEEDLEFT HORIZONTALLY, FOLLOWINGSHAPE OF CURVEDLINES, TO CROSSWINDANGLE FROM TAKEOFFHEADING,POINT B. PROCEEDVERTICALLYUPWARDTO CROSSWIND COMPONENTLIMITSCALE AND READCROSSWINDCOMPONENT............................. (NOTE: CHECK SECTION1A CROSSWINDLIMITATIONS) PROCEEDLEFT HORIZONTALLY TO HEADWINDCOMPONENT SCALE AND READUNFACTOREDHEADWINDCOMPONENT................................... (NOTE: NO HEADWINDBENIFIT)

Figure 1C-2. Headwind and Crosswind Component Chart

170° ° . 200 °

3.........30

20 KNOTS

10 KNOTS

17.5 KNOTS

412FS62.3-1C-22

T N I Section2C R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O & 62.4 BHT-412-FMS-62.3

FAA APPROVED

PARTC--RUNWAY

2C-2. PREFLIGHTPLANNING

All Category A takeoffs and landings should begin with thorough planning. Each segment, 1, 2, or 3; should be planned for by utilizing correct charts contained and identified within Section 4A. Each segment can be planned and actual climb gradient calculatedto ensure necessary obstacle clearanceis available. Values derived from performance charts should be readily available to flying pilot (FP) and reviewed, as necessary, just prior to takeoff.

2C-7. BEFORETAKEOFF

2C-7-A. POWER ASSURANCE CHECK

check shall be Power assurance prior to accomplished satisfactorily conducting Category A operations (See - POWERPLANT). LIMITATIONS

2C-7-B. OEI (TARGET) TORQUE

For CategoryA Runwayoperations(takeoffor landing), refer to OEI (target)torque on Power Available2.5 Minute RatingChart(Section4C) to obtain OEI (target) torque available at time of operation.

2C-7-C. RUNWAY LENGTH REQUIRED

Refer to Single Engine Rejected Takeoff Distance Chart (Section 4C) for required length.

2C-8. TAKEOFF/LANDING

2C-8-A. GENERAL INSTRUCTIONS For CategoryA Runwayoperations,Takeoff DecisionPoint (TDP)is a timed point in takeoff profile (Figure2C-1).For TDP determinations,timing starts at commencementof forward cyclic input for a level acceleration.This time is a function of altitude and temperatureand is depictedon Takeoffdecision point chart (Section4C).

Actual time count for TDP can be determined from instrument panel clock, or can be counted by pilot (i.e. one thousand one, one thousand two, etc.). At TDP, any remaining power (up to and including takeoff power) is applied and accelerationis continued to Vy (70 KIAS) and 200feet (61 meters) AGL. After 200 feet AGL (61 meters), Np is adjusted to 100% prior to accelerationpast 80 KIAS.

73


FAA APPROVED

2C-8-B. TAKEOFF

After obtaining Vy and 200 ft above runway, adjust Np to 100% prior to accelerating past 80 KIAS.

PLANNING CONSIDERATIONS:

Review any factors which may affect Segment 1 (Takeoff). Verify wind direction and velocity. Verify correct gross weight for takeoff. Determine 21/2minute power (Target Torque). Determine TDP (Time in seconds).

Seat(s) - adjusted.

PLANNING CONSIDERATIONS:

Review any factors which may affect landing.

Verify wind direction and velocity. Verify correct gross weight for landing.

Flight controls - adjust friction.

Altimeter - Set.

Throttles - full open.

Set to 15 ft.

RPM INCR/DECR switch airspeed below 80KIAS.

ALTN STATIC SOURCE - PRI.

RPM INCR/DECR switch - 103% Np Other instruments ranges. Runway -

LANDING

Seat(s) - adjusted.

Collective - Full down (flat pitch).

Radar altimeter -

2C-8-C.

-

Normal

operating

Clear.

Other instruments ranges.

-

-

103% Np at

Normal

operation

Altimeter - Set.

Radar Altimeter - Set (LDP) 50 ft.

Hover - 10 ft (Figure 2C-1).

ALTN STATIC SOURCE -

Mast Torque - Note.

Conduct normal approach (Figure 2C-1) to arrive at 50 ft and 40 KIAS. From LDP, a slow deceleration to 4 ft hover on runway terminates approach.

Collective/Cyclic - Apply forward cyclic to approximately 10 ° nose down while simultaneously applying an additional 10% mast torque.

A level acceleration is begun at same time countdown to TDP is begun.

At TDP, begin a positive nose up cyclic input while applying remaining takeoff power (5%) and accelerating to Vy (70 KIAS).

74

PRI.

I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O RUNWAY

Normal Takeoff/Landing Profile

-

TAKEOFF PROFILE LANDINGPROFILE

T N Section3C I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-62.3 & 62.4

3C-3. ENGINE

For continuedtakeoff, proceed as follows:

3C-3-A. SINGLE ENGINEFAILURE (MALFUNCTIONS

Immediately rotate nose to VTOSS climb attitude while simultaneously applying collective to predeterminedOEItorque.

RESULTING IN LOSS OF POWER: I.E. ENGINE FAILURE, LOW SIDE GOVERNOR FAILURE)

3C-3-A-1. DURING TAKEOFF PRIOR TO TDP (REJECTED TAKEOFF) Refer to Figure 3C-1 Emergency Landing/ TakeoffProfile.

An enginefailure prior to reachingTDPabove runway will necessitate a sliding landing to runway. Upon engine failure, pilot will level skids and allow helicopter to settle back to runway. Maximum OEI power may not be necessary to complete landing but may be utilized, if required. Touchdown on runway should be slightly nose high to minimize slide. Add collective, as required, to cushion landing and apply cyclic inputs to eliminatelateral drift. Performnormal shutdown.

3C-3-A-2. DURINGTAKEOFF AT OR AFTER TDP (COMPLETEDTAKEOFF)

Refer to Figure 3C-1 Emergency Landing/ Takeoff Profile In event of an engine failure during takeoff at TDP.

Maintain21/2minute OEIpower (TargetTorque) and VTOSS (40 KIAS)while climbing to 200feet (61 meters)above runway.Upon reaching 200 feet (61 meters) above runway, accelerate to Vy (70 KIAS). At end of 2 1/2 minute OEI limit, decrease power to 30 minute or MCPpower setting and 100%/ Np / NR . At safe altitude, apply OEI procedures (Section 3, SINGLE ENGINE FAILURE-INFLIGHT, Basic Flight Manual). 3C-3-A-3. DURING LANDING PRIOR TO OR AT LDP (BALKED LANDING)

Refer to Figure 3C-1 Emergency Landing/ Takeoff Profile in event of enginefailure prior to LDP.

Upon engine failure, pilot shall apply collectiveto predeterminedOEI enginetorque (target torque) while accelerating to Vy (70 KIAS).

At end of 2 1/2 minute OEI limit, decrease power to 30 minute or MCPpower setting and 100%/ NP / NR. At safe altitude, apply OEI procedures (Section 3, SINGLE ENGINE FAILURE,Basic Flight Manual). 3C-3-A-4. DURING LANDING PRIOR TO OR AFTER LDP (COMPLETEDLANDING)

Refer to Figure 3C-1 Emergency Landing/ TakeoffProfile. An engine failure after reaching LDP above runway will necessitatea landing to runway.


BHT-412-FMS-62.3 & 62.4

Upon engine failure, pilot will set predetermined 2 1/2 minute OEI power (target torque) and allow helicopter to settle to runway. Touchdown on runway should be slightly nose high to minimize slide. Add collective as required to cushion landing and apply cyclic inputs to eliminate lateral drift.

Before TDP:

Perform normal engine shutdown.

After TDP:

3C-3-A-5. WARNING/CAUTIONLIGHTS (ANY MALFUNCTIONTHAT DOES NOT RESULT IN ENGINE POWER LOSS)


If Master Caution light illuminates, continue takeoff profile. After obtaining Vy and 200 feet altitude, take corrective action (see Basic Flight Manual).

At or after LDP:

Control rotor RPM.


77

I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O RUNWAY

Emergency Landing/Takeoff Profile

---

-

LDP

- -

TAKEOFF/REJECT

-

- -

LANDING

Vy (70 KIAS) 200 FEET

T N I Section4C R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4

FAA APPROVED

PART C - RUNWAY

4C-1. INTRODUCTION Performance supplement manufacturer's with installation

in this data presented on engine are based minimum specification power losses.

4C-2. POWER ASSURANCE CHECK

Refer to power assurance check chart in Basic Flight Manual Section 4, PERFORMANCE, or appropriate FMS for installed engines (see LIMITATIONS - POWER PLANT). Instructions for chart use can be found on chart.

Power assurance check chart is based on a with engine specification minimum losses and proper rigging of installation engine controls. If engine performance does not meet that shown in power assurance check, Category A operations will not be attempted and steps must be taken to ascertain causes of engine power loss.

4C-6. TAKEOFF DISTANCE/ PROFILE

For runway operations, takeoff distance/ profile is horizontal distance beginning at point of takeoff hover and progressing to a point 1000 feet (305 meters) above takeoff surface. This distance/path, after engine failure, can be defined in three segments.


WAT limits (Figure 4C-1).

Headwind and crosswind component (Figure 4C-2).

2 1/2minute power rating (Figure 4C-3). Single engine rejected takeoff distance (Figure 4C-4). Takeoff decision point (Figure 4C-5).

Point of takeoff to a point 35 feet (10.7 meters) above takeoff surface and VToss/positive rate of climb (Figure 4C-6).

Segment 2 (VToss climb/gradient):

Distance/path from VTOSSand 35 feet (10.7 meters) to 200 feet (61 meters) above takeoff surface (Figure 4C-7). Distance for level acceleration from VTOSSto Vy (Figure 4C-8).

Segment 3 (Vy climb/gradient):

Distance/path from 200 feet (61 meters) above to 1000 feet (305 meters) above takeoff surface (Figure 4C-9).

EXAMPLES, RUNWAY:

What is takeoff distance/path, given following conditions? Engines - PT6T-3D. Altitude - 1500feet (457 meters) HP. Temperature - 5°C. Wind - 10 knots on helicopter nose.

79


Step 1. Maximum allowable gross weight: WAT chart (Figure 4C-1) -11,900 pounds.

Crosswind component (Figure 4C-2) 0 knots.

Step 2.

OEI (Target) Torque:

FAA APPROVED

ANSWER:

Total Takeoff Distance Step 3 -

+Step4

3,100

+Step 5 -

16,000

2 1/2 minute rating (Figure 4C-3) - 67% torque.

Step 3.

Takeoff decision point (TDP) (Figure 4C-5)- 5 seconds.

Takeoff distance to VTOSSand 35 feet (10.7 meters) (Figure 4C-6) -1,200 feet (366 meters). VTOSSclimb (Segment 2):

Distance from VTOSS/35feet (10.7 meters) to 200 feet (61 meters) (Figure 4C-7). 11% climb gradient.

165 feet (50.3 meters) +.11- 1,500 feet (457 meters). Distance for level acceleration from VTOSSto Vy (Figure 4C-8) - 1,600 feet (487.7 meters).

Step 5.

Vy Climb (Segment 3:

Distance from 200 feet (61 meters) to 1,000 feet (305 meters) above takeoff surface (Figure 4C-5).

5% climb gradient. 800 feet (244 meters) +.05 - 16,000 feet (4877meters).

80

20,300 Feet (6,187 Meters)

Takeoff (Segment 1):

Rejected takeoff distance (Figure 4C-4) - 625 feet (190.5 meters).

Step 4.

1,200

-

4C-7. CLIMB AND DESCENT

4C-7-A. SINGLE ENGINE RATE OF CLIMB CHARTS These charts and their locations are provided for additional Information. Vy (70 KIAS):

See Basic Flight Manual or appropriate FMS for installed engines (see LIMITATIONS- POWER PLANT).

VTOSS(40 KIAS):

Rate of climb chart (VToss - 40 KIAS) (Figure 4C-11).

This chart is a function of gross weight, Hp OAT, and is same for PT6T3D, DE, DF engines at 2 1/2 minute power.

4C-9. LANDING DISTANCE

Single Engine Landing Distance chart (Figure 4C-10) provides landing distances required to clear a 50 feet (15 meter) obstacle for stated outside air temperature, pressure altitudes, and appropriate Weight-Altitude-Temperature chart. Landing distances are based on an approach condition of 40 KIAS (VToss), zero wind, and less than 500 FPM rate of descent.


FAA APPROVED

LIMITS WEIGHT-ALTITUDE-TEMPERATURE RUNWAYTAKEOFF AND LANDING

PT6T-3DENGINE 103%RPM(AEO) GENERATOR150AMPS

36

HEATEROFF

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

52

54

GROSS WEIGHT - LB

412FS62.3-4C-1-1



FAA APPROVED

WEIGHT-ALTITUDE-TEMPERATURE LIMITS RUNWAYTAKEOFF AND LANDING

PT6T-3DEENGINE 103%RPM(AEO) GENERATOR 150AMPS

HEATEROFF

36

38

GROSS WEIGHT -100 KG 40 42 44 46 48 50

52

54

412F562.3-4C-1-2



LIMITS WEIGHT-ALTITUDE-TEMPERATURE RUNWAY TAKEOFF AND LANDING

HEATER OFF


36 6000 '

38 '

1.8

5000 --

GROSS WEIGHT- LB

412FS62.3-4C-1-3



1. 2. 3. 4. 5. 6. 7. 8.

EXAMPLE: TAKEOFFHEADING..................................................................... REPORTEDWIND DIRECTION............................................................ DIFFERENCEBETWEENWIND DIRECTIONAND TAKEOFFHEADING........................... REPORTEDWIND SPEED ............................................................... ENTER CHARTAT REPORTEDWIND SPEED, POINTA. PROCEEDLEFT HORIZONTALLY, FOLLOWINGSHAPE OF CURVEDLINES, TO CROSSWINDANGLE FROMTAKEOFFHEADING,POINT B. PROCEEDVERTICALLYUPWARDTO CROSSWIND COMPONENTLIMITSCALE AND READCROSSWINDCOMPONENT............................. (NOTE: CHECK SECTION1A CROSSWINDLIMITATIONS) PROCEEDLEFT HORIZONTALLY TO HEADWINDCOMPONENT SCALE AND READ UNFACTOREDHEADWINDCOMPONENT................................... (NOTE: NO HEADWINDBENIFIT)

Figure 4C-2. Headwind and Crosswind Component

84

FAA APPROVED

°

170 170° ° 0 10 KNOTS

0 KNOTS

10 KNOTS 412FS62.3-4C-2


FAA APPROVED

POWER AVAILABLE 2 1/2MINUTE RATING

HEATER OFF


6000

5000

ENGINE (TARGET)

TORQUE - %Q

Figure 4C-3 Power Available 2 1/2 Minute Rating 85


FAA APPROVED

SINGLE ENGINE REJECTEDTAKEOFF DISTANCE RUNWAY

PT6T-3D/DE/DF ENGINE 100%RPM(OEI) GENERATOR 150AMPS

HEATEROFF

*MINIMUM DISTANCE 625 FEET

ENGINE FAILURE BEFORE TDP

TDP

REJECT TAKEOFF DISTANCE

412FS62.3-4C-4

Figure 4C-4 Single Engine Rejected Takeoff Distance


FAA APPROVED

TAKEOFF DECISION POINT (TDP) RUNWAY

HEATER OFF

PT6T-3D/DE/DF ENGINE 103% RPM (AEO) GENERATOR 150 AMPS

6000 FT. H D

-40

-20

0

20

40

60 0

OAT- °C

12

3

8

4 5 6 7 TDP-SECS

9 10

NORMALTAKEOFF

Vy(70 KIAS)

TDP

VTOSS

10 FT

SECONDS TAKEOFF DECISIONPOINT

412FS62.3-4C-5

Figure 4C-5 Takeoff Decision Point (TDP) 87


FAA APPROVED

SINGLE ENGINE TAKEOFFDISTANCE TO 35 FT ABOVE TAKEOFFSURFACE RUNWAY


HEATER OFF

2000

* MINIMUM DISTANCE 1150 FEET

ENGINEFAILUREAFTER TDP

35 FT

TDP

VTOSS RATEOF CLIMB POSITIVE

VROT

10 FT

OEITAKEOFFDISTANCE

412FS62.3-4C-6

Figure 4C-6 Single Engine Takeoff Distance To 35 (10.7 Meters) Feet Above Takeoff Surface 88

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TAKEOFF FLIGHT PATH FROM VTOSS TO 200 FEET ABOVE TAKEOFFSURFACE

PT6T-3D/DE/DF ENGINE 2.5 MIN POWER 100% RPM (OEI) GENERATOR 150 AMPS

VTOSS= 40 KIAS

HEATER OFF

6000

5000

z

0

o

Figure 4C-7. Takeoff Flight Path 35 Feet (10.7 Meters) To 200 Feet (61 Meters) AGL 89


BHT-412-FMS-62.3 & 62.4

HORIZONTALDISTANCEFORLEVELACCELERATION

PT6T-3DENGINE 30 MINUTEPOWER 100%RPM(OEI) GENERATOR 150AMPS

FROMVTOSSTO VY ZEROWIND

HEATEROFF

HORIZONTAL DISTANCE - METERS 150 200 250 300 350 400 450 500 550 600 650 700 750 800 850

412FS62.3-4C-8-1

Figure 4C-8 Acceleration VTOSSTo Vy (Sheet 1 of 3) 90


FAA APPROVED

HORIZONTALDISTANCEFORLEVELACCELERATION TOVY VTOSS ~FROM ZERO WIND

PT6T-3DEENGINE 30 MINUTEPOWER 100%RPM(OEI)

HEATEROFF

HORIZONTALDISTANCE- METERS

12000

400

600

800

1000

Figure 4C-8

1200 1400 1600 1800 2000 2200 HORIZONTAL DISTANCE - FEET

Acceleration

VTOSS To Vy (Sheet

2400

2600

2800

2 of 3)

91


BHT-412-FMS-62.3& 62.4


PT6T-3DF ENGINE 30 MINUTEPOWER 100%RPM(OEI) 150AMPS GENERATOR

I-


HEATEROFF

HORIZONTAL DISTANCE - METERS 150 200 250 300 350 400 450 500 550 600 650 700 750 800

400

600

800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 HORIZONTAL DISTANCE - FEET

412FS62.3-4C-8-3

Figure 4C-8 Acceleration VTOSSTo Vy (Sheet 3 of 3)


FAA APPROVED

TAKEOFFFLIGHTPATH200 TO 1000 FEET ABOVETAKEOFF SURFACE


PT6T-3DENGINE POWER MAXCONTINUOUS 100% RPM(OEI) 150 AMPS GENERATOR 6000

5000

1000

z

CLIMB GRADIENT -%

412FS62.3-4C-9-1

Figure 4C-9. Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 1 of 3) 93


FAA APPROVED

TAKEOFFFLIGHT PATH200 TO 1000 FEET ABOVE TAKEOFF SURFACE

PT6T-3DEENGINE 30 MINUTEPOWER 100%RPM(OEI) GENERATOR 150 AMPS

VY =70 KIAS

HEATEROFF

CLIMBGRADIENT-

Figure 4C-9 Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 2 of 3)


FAA APPROVED

TAKEOFF FLIGHT PATH200 TO 1000 FEET ABOVE TAKEOFF SURFACE

ENGINE PT6T-3DF 30 MINUTEPOWER 100%RPM(OEI) GENERATOR150AMPS

Figure 4C-9


Takeoff Flight Path 200 Feet (61 Meters) To 1000 Feet (305 Meters) (Sheet 3 of 3) 95


FAA APPROVED

SINGLE ENGINE LANDING DISTANCE RUNWAY

HEATEROFF

PT6T-3D/DE/DF ENGINE 100%RPM(OEI) GENERATOR 150AMPS

6000 FT. HD

-40

-20

0

20

40

60 0

OAT- °C

200

400

600

800

1000

LANDING DISTANCE - FT

ENGINEFAILUREBEFOREOR AFTER LDP

LDP

VTOSS (40 KIAS)

LESSTHAN500 FPM RATE OF DESCENT

OEI LANDINGDISTANCE

412FS62.3-4C-10

Figure 4C-10 Single Engine Landing Distance


FAA APPROVED

SINGLE ENGINERATE OF CLIMB

PT6T-3D/DEIDFENGINE 2.5 MINUTE 100%RPM (OEI) GENERATOR150 AMPS

VTOSS = 40 KIAS HEATER OFF

RATEOF CLIMB - M/SEC

-4

-3

L

400

600

800 1000 1200 1400 1600 1800 2000 2200

RATEOF CLIMB - FT/MIN

412FS62.3-4C-11

Figure 4C-11. Single Engine Rate of Climb (VTOSS- 40KIAS) 97/98


4-1.

FUEL FLOW VS AIRSPEED

Fuel flow charts (Figure 4-1) are based on estimates and limited flight test data. Fuel consumption may vary betweenengines. It is recommended that operators conduct measurementsto be used to adjust presented data as required. This data does not include effects of bleed air on fuel consumption. Fuel flow is applicable to basic helicopter without any optional equipment which would appreciably affect lift, drag, or power available.

BHT-412-FMS-62.3 & 62.4

4-2. SINGLE ENGINEHOVER

Single engine maximum gross weight for hover chart (Figure 4-2)is based upon engine manufacturer's minimum specification power for PT6T-3D/DE/DF engines.

Rev. 1

99

T FUEL FLOW VS AIRSPEED N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4

MANUFACTURER'S DATA

TWIN ENGINE OPERATION (100% RPM) ZEROWIND

PRESSURE ALTITUDE = SEA LEVEL

3

20

130

135

Figure 4-1

100

Fuel flow at 100% RPM (Sheet 1 of 18)

140

T FUEL FLOW VS IAIRSPEED N R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O TWIN ENGINE OPERATION

(100% RPM)

ZERO WIND

PRESSURE ALTITUDE = SEA LEVEL OAT

+15 ° C


Figure 4-1 Fuel flow at 100% RPM (Sheet 2 of 18)

412FS62.3-MD-4-1-2

T N FUEL FLOW VS AIRSPEED I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4

MANUFACTURER'S DATA


(100% RPM)

ZEROWIND

PRESSURE ALTITUDE = SEA LEVEL

CLEAN CONFIGURATION

0

4

60

70

80

90

100

TRUE AIRSPEED -

60

70

80

90

100

110

KNOTS

110

1


Figure 4-1


412FS62.3 MD-4-1-3


m

130

130


140

140

T FlUEL FLOW VS AIRSPEED N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

BHT-412-FMS-62.3 & 62.4

TWIN ENGINEOPERATION (100% RPM) ZEROWIND

PRESSURE ALTITUDE = 2000 FEET OAT = +11 °C FUEL FLOW INCREASE FOR POP-OUT FLOATS

8

7

CLEAN

CONFIGURATION

z

130

140

412FS62.3-MD-4-1-5

Figure 4-1

104


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O INDICATED AIRSPEED -KNOTS Figure 4-1


105

T FUEL FLOW VS AIRSPEED N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3& 62.4

MANUFACTURER'S DATA


(100% RPM)

ZEROWIND

PRESSURE ALTITUDE = 4000 FEET 20

FUEL FLOW INCREASE FOR POP-OUT FLOATS I

10

Figure 4-1


T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FUEL FLOW VS AIRSPEED TWIN ENGINEOPERATION (100% RPM) ZEROWIND

PRESSURE ALTITUDE = 4000 FEET

Figure 4-1


107


MANUFACTURER'S

DATA

FUEL FLOW VS AIRSPEED TWIN ENGINE OPERATION

I

(100% RPM)

ZEROWIND

PRESSURE ALTITUDE = 4000 FEET OAT = +27°C FUEL FLOW INCREASE FOR POP-OUT FLOATS

10

60

70

80

90

TRUE AIRSPEED

100

1 10

120

130

140

- KNOTS

7

412FS62.3-MD-4-1-9

Figure 4-1



130

412FS62.3-MD-4-1-10

Figure 4-1


109


MANUFACTURER'S DATA

FUEL FLOW VS AIRSPEED TWIN ENGINE OPERATION

(100% RPM)

ZERO WIND

PRESSURE ALTITUDE = 6000 FEET OAT = + 3°C

20

FUEL FLOW INCREASE FOR POP-OUT FLOATS

10

60

70

80

90

100

110

1

TRUE AIRSPEED - KNOTS

60

70

80

55

65

75

90 100 TRUE AIRSPEED 85

INDICATED

Figure 4-1

110 KNOTS

95

AIRSPEED

120

105

-

KNOTS


115

130

140

125

412FS62.3-MD-4-1.11

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Figure 4-1


111


MANUFACTURER'S

DATA



OAT= -21°C

FUEL FLOW INCREASE FOR POP-OUT FLOATS

8

140

120

130

412FS62.3 MD-4-1-13

Figure 4-1



140

115

125

INDICATED AIRSPEED - KNOTS Figure 4-1


113


MANUFACTURER'S DATA


PRESSURE ALTITUDE = 8000 FEET OAT = +19°C

m

IFUEL FLOW INCREASE FOR POP-OUT FLOATS


Figure 4-1


T FUEL FLOW VS AIRSPEED N I R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-62.3 & 62.4

MANUFACTURER'S DATA


ZERO WIND

(100% RPM)

PRESSURE ALTITUDE = 10000

FEET

OAT = -25°C

20

FUEL FLOWINCREASE FOR POP-OUTFLOATS

10

--

60

70

130

120

110

100

90

80

140


CLEANCONFIGURATION

XMSN LIM

GW 1000

60

55

70

65

LB

110 90 100 TRUEAIRSPEED - KNOTS

80

75

85

95

130

120

105

INDICATEDAIRSPEED-KNOTS

115

140

125

412FS62.3-MD-4-1-16

Figure 4-1 Fuel flow at 100% RPM (Sheet 16 of 18)

115


MANUFACTURER'SDATA


(100% RPM)

ZERO WIND

PRESSURE ALTITUDE = 10000

OAT=-

FEET

5°C

FUEL FLOWINCREASEFOR POP-OUTFLOATS

10

60

70

80

90

100

120

110

130

140


CLEANCONFIGURATION

XMSN LIM

5-

I

MAXEND

60

50

70

60

80 70

90 100 110 TRUEAIRSPEED- KNOTS 80

90

120

100

INDICATED AIRSPEED - KNOTS Figure 4-1 Fuel flow at 100% RPM (Sheet 17 of 18)

116

130

110

140

120

412FS62.3-MD-4-1-17

T FUEL FLOW VS AIRSPEED N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

BHT-412-FMS-62.3 & 62.4




MANUFACTURER'S DATA

& 62.4

SINGLE ENGINE HOVER CEILING OUT OF GROUND EFFECT

PT6T-3D/DE/DF ENGINE 2.5 MIN POWER 103 % RPM (OEI) GENERATOR 150 AMPS

SKID HEIGHT60 FT (18.3METER) HEATEROFF

GROSS WEIGHT - KG

3200

3600

4000

4400

4800

6000

5000

8000

9000

10000

11000

GROSS WEIGHT - LB

Figure 4-2. Single engine hover ceiling OGE

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O This supplement shall be attached to Model 412 or 412EP Flight Manual when SELF SEALINGFUEL CELLS are installed.

Information contained herein supplements Information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, consult basic Flight Manual.

COPYRIGHT NOTICE COPYRIGHT 1997 BELL HELICOPTER TEXTRON INC. ANDBELLHELICOPTER TEXTRON INC. A DIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

19 SEPTEMBER 1997 REVISION 1 - 22 OCTOBER1997

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-63.2& 63.3 & 63.4

NOTICEPAGE





LOGOFREVISIONS

Original ........... Revision .........

0 .............. 1 ..............

19 SEP 97 22 OCT 97

LOGOF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

REVISION NO.

1 - 2 .....................................

T itle .................................... NP ......................................

1 0

A .......................................

1

B. ...................................... 1/11 .......................................

PAGE

0 0

0

3- 4 ........... .. ................... 4A/4B ................................... 5- 7 ..................................... 8- 10 .................................... 11 30 .................................. 31/32 ...................................

1 1 1 0 1 1

NOTE


A

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-63.2 & 63.3 & 63.4


Original...........0..............19

SEP 97

APPROVED:

MANAGER/


DATE:


GENERAL INFORMATION

The customizing-lower self sealing fuel cell installation minimizes the loss of fuel in the event of puncture damage by small objects.

Due to the increased wall thickness of the cell, the total and useable fuel capacities are less than those of the basic helicopter.

T N I * Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED


BHT-412-FMS-63.2

to retain gross necessary, within allowable limits.

& 63.3 & 63.4

weight

CG

Actual weight change shall be determined after installation and ballast readjusted, if

1

T N I Section2 R P N O R I O T F C U T D O Section 3 O N R E P E E R R C S Section4 NG O N I R W O E I F V NLY O BHT-412-FMS-63.2& 63.3 & 63.4




FAA APPROVED

BHT-412-FMS-63.2& 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-7 - Provides lateral CG data for right side installation.

5-7. FUEL LOADING

Fuel loading tables lists usable fuel quantities in 10 gallon (40 liter) increments, with weights and moments in both english and metric units for balance computation. Critical fuel loading for computing most forward and aft CGs are denoted.

5-7-A. BASICSYSTEM- SELF SEALINGFUEL CELL.

Total capacity: 328.3 U.S. gallons (1242.2 liters). Usable fuel: 321.3 U.S. gallons (1215.7 liters). * *

Tables 5-1 and 5-2 - Provides longitundinal CG data for approved fuels. Tables 5-3 and 5-4 - Provides lateral CG data for approved fuels.

5-7-C. BASIC SYSTEMWITH BOTH LONGRANGEAUXILIARYFUEL TANKS. Refer to BHT-412-FMS-17.2/17.3/17.4.


Table 5-8 - Provides longitundinal CG data for approved fuels. Table 5-9 - Provides lateral CG data for approved fuels.

5-7-D. BASIC SYSTEMWITHONE SEATTYPEAUXILIARYFUELTANK (LEFTOR RIGHT). Refer to BHT-412-FMS-25.2/25.3/25.4.

5-7-B. BASICSYSTEMWITHONE LONGRANGEAUXILIARYFUEL TANK (LEFTOR RIGHT). Refer to BHT-412-FMS-17.2/17.3/17.4.


Total capacity: 344.6 U.S. gallons (1304.2 liters). Usable fuel: 337.6 U.S. gallons (1277.7 liters). * * *

Table 5-10 - Provides longitundinal CG data for approved fuels. Table 5-11 - Provides lateral CG data for left side installation. Table 5-12 - Provides lateral CG data for right side installation.

Table 5-5 - Provides longitundinal CG data for approved fuels. Table 5-6 - Provides lateral CG data for left side installation. Rev.1

3

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 5-7-E. BASICSYSTEMWITHBOTH SEATTYPEAUXILIARYTANKS. Refer to BHT-412-FMS-25.2/25.3/25.4.

Total capacity: 360.9 U.S. gallons (1365.9 liters).

Rev.1

Usable fuel: 353.9 U.S. gallons (1339.4 liters). *

Table 5-13 - Provides longitundinal CG data for approved fuels. * Table 5-14 - Provides lateral CG data for approved fuels.

BHT-412-FMS-63.2 & 63.3 & 63.4


U.S. GALLONS

Usable fuel loading table self-sealing tanks (English) Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal) WEIGHT (lb)

LONGCG (in.)

MOMENT (in-lb)

139.5 9486 68 10 18999 139.7 136 20 28519 139.8 30 204 139.9 38053 272 40 47566 139.9 340 50 51903 139.9 371 54.6 58589 408 143.6 60 70448 148.0 476 70 151.4 82362 544 80 94309 612 154.1 90 106352 680 156.4 100 158.2 118334 748 110 130234 159.6 816 120 142147 884 160.8 130 161.5 149711 927 136.3 151939 159.6 952 140 158100 155.0 150 1020 151.0 164288 1088 160 169334 1138 148.8 167.3 172475 149.2 1156 170 150.7 184457 1224 180 196384 152.0 1292 190 208352 1360 153.2 200 220483 154.4 1428 210 232329 1496 155.3 220 243984 1564 156.0 230 156.7 252914 1614 237.3 254918 1632 156.2 240 263160 1700 154.8 250 153.0 270504 1768 260 278338 1836 151.6 270 285981 1904 150.2 280 149.5 290628 1944 285.9 295406 149.8 1972 290 307428 150.7 2040 300 151.5 319362 2108 310 331405 152.3 2176 320 332994 152.4 321.3 2185 * Most critical amount for most forward C.G.condition at weight empty is 6580pounds or greater. ** Most critical fuel amount for most aft C.G. condition at weight empty is up to 6450pounds. *** Most critical fuel amount for most aft C.G.condition at weight empty Is 6450pounds or greater. Weightsgiven are nominal weights at 15 C. NOTE This table is invalid with auxiliary fuel tank(s) installed. 412-FMS-63-5-1-

Rev. 1

4A/4B

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-1M. Usable fuel loading table self-sealing tanks (Metric) Jet A, JP-5, JP-8 (0.815 kg/I)

LITERS

(kg)

(mm)

(kg-mm/100)

40 80 120 160 200 206.7 240 280 320 360 400 440 480 515.7 520 560 600 633.2 640 680 720 760 800 840 880

32.6 65.2 97.8 130.4 163.0 168.5 195.6 228.2 260.8 293.4 326.0 358.6 391.2 420.3 423.8 456.4 489.0 516.1 521.6 554.2 586.8 619.4 652.0 684.6 717.2

3551 3551 3551 3553 3553 3553 3691 3795 3876 3945 3998 4041 4074 4102 4087 3962 3851 3780 3787 3825 3863 3894 3922 3950 3973

1158 2315 3473 4633 5791 5987 7220 8660 10109 11575 13033 14491 15937 17241 17321 18083 18831 19509 19753 21198 22668 24119 25571 27042 28494

898.2

732.0

3980

29134

920 960 1000 1040 1080 1082.1 1120 1160 1200 1215.7

749.8 782.4 815.0 847.6 880.2 881.9 912.8 945.4 978.0 990.8

3955 3912 3871 3835 3797 3797 3820 3840 3861 3871

29655 30607 31549 32505 33421 33486 34869 36303 37761 38354

* Most critical amountfor most forward C.G.conditionat weight emptyis 2984kilogramsor greater. ** Most critical fuel amount for most aft C.G. condition weight empty Is up to 2925kilograms. ***Most criticalfuel amount for most aft C.G.conditionat weightemptyis 2925kilogramsor greater. Weightsgiven are nominal weights at 15 C. NOTE This table is invalid with auxiliary fuel tank(s) installed.

Rev. 1

5


Table 5-2. Usable fuel loading table self-sealing tanks (English) Jet B, JP-4 (6.5 Lbs/Gal)

GALLONS

*

(lb)

(in.)

(in-lb)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 136.3 140 150 160 167.3

65 130 195 260 325 355 390 455 520 585 650 715 780 845 886 910 975 1040 1087

139.5 139.7 139.8 139.9 139.9 139.9 143.6 148.0 151.4 154.1 156.4 158.2 159.6 160.8 161.5 159.6 155.0 151.0 148.8

9068 18161 27261 36374 45468 49665 56004 67340 78728 90149 101660 113113 124488 135876 143089 145236 151125 157040 161746

170 180 190 200 210 220 230 237.3

1105 1170 1235 1300 1365 1430 1495 1542

149.2 150.7 152.0 153.2 154.4 155.3 156.0 156.7

164866 176319 187720 199160 210756 222079 233220 241631

240 250 260 270 280 285.9 290 300 310 320 321.3

1560 1625 1690 1755 1820 1858 1885 1950 2015 2080 2088

156.2 154.8 153.0 151.6 150.2 149.5 149.8 150.7 151.5 152.3 152.4

243672 251550 258570 266058 273364 277771 282373 293865 305273 316784 318211

* Most critical amount for most forward C.G. condition at weight empty is 6580pounds or greater. * Most critical fuel amount for most aft C.G. condition at weight empty is up to 6450pounds. Most critical fuel amount for most aft C.G.condition at weight empty is 6450pounds or greater. Weightsgiven are nominal weights at 15 C. NOTE This table is invalid with auxiliary fuel tank(s) installed. 412-FMS63-5-2-1

6

Rev.1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-2M. Usable fuel loading table self-sealing tanks (Metric) Jet B, JP-4 (0.779 kg/I)

LITERS

*

*

(kg)

(mm)

(kg-mm/100)

40 80 120 160 200 206.7 240 280 320 360 400 440 480 515.7 520 560 600 633.2 640 680 720 760 800 840 880

31.2 62.3 93.5 124.6 155.8 161.0 187.0 218.1 249.3 280.4 311.6 342.8 373.9 401.7 405.1 436.2 467.4 493.3 498.6 529.7 560.9 592.0 623.2 654.4 685.5

3551 3551 3551 3553 3553 3553 3691 3795 3876 3945 3998 4041 4074 4102 4087 3962 3851 3780 3787 3825 3863 3894 3922 3950 3973

1108 2212 3320 4427 5536 5720 6902 8277 9663 11062 12458 13853 15233 16478 16556 17282 18000 18647 18882 20261 21668 23052 24442 25849 27235

898.2

699.7

3980

27848

920 960 1000 1040 1080 1082.1 1120 1160 1200 1215.7

716.7 747.8 779.0 810.2 841.3 843.0 872.5 903.6 934.8 947.0

3955 3912 3871 3835 3797 3797 3820 3840 3861 3871

28345 29254 30155 31071 31944 32009 33330 34698 36093 36658

* Most critical amountfor most forward C.G.conditionat weightemptyis 2984kilogramsor greater. ** Most critical fuel amount for most aft C.G. condition at weight empty is up to 2925kilograms. *** Most critical fuel amountfor most aft C.G.conditionat weightemptyis 2925kilogramsor greater. Weights given are nominal weights at 15C. NOTE This table is invalid with auxiliary fuel tank(s) installed. 412-FMS43-5-2-2

Rev. 1

7

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-63.2 & 63.3 & 63.4 Table 5-3. U.S.

GALLON

10 20

30 40

50 54.6 60 70 80 90 100 110 120 130 136.3 140 150 160 167.3 170 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 321.3

Lateral usable fuel loading table self-sealing tanks (English) Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal) WEIGHT (Lbs.)

LAT CG (in.)

0 0 0 0 0 0 -0.06 -0.05 -0.05 -0.04 -0.04 -0.04 -0.03 -0.03 -0.03 -0.31 -0.50 -0.63 -0.70 -0.69 -0.65 -0.61 -0.58 -0.55 -0.53 -0.51 -0.49 -0.47 -0.45 -0.43 -0.41 -0.40 -0.39 -0.37 -0.36 -0.36

68 136 204 272 340 371 408 476 544 612 680 748 816 884 927 952 1020 1088 1138 1156 1224 1292 1360 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2108 2176 2185

MOMENT (in-lb)

0 0 0 0 0 0 -24 -24 -27 -24 -27 -30 -24 -27 -28 -295 -510 -685 -796 -798 -796 -788 -789 -785 -793 -798 -800 -799 -796 -789 -781 -789 -796 -780 -783 -787

* Most critical fuel amount for left side most lateral C.G.condition. Weights given are nominal weights at 15 C. NOTE


412-FMS-63-5-3-1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-3M.

LITERS

40 80 120 160 200 206.6 240 280 320 360 400 440 480 515.6 520 560 600 633.0 640 680 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1215.7

Lateral usable fuel loading table self-sealing tanks (Metric) Jet A, A-1, JP-5, JP-8 (0.815 kg/I) WEIGHT (kg)

LAT CG (mm)

32.6 65.2 97.8 130.4 163.0 168.4 195.6 228.2 260.8 293.4 326.0 358.6 391.2 420.2 423.8 456.4 489.0 515.9 521.6 554.2 586.8 619.4 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 912.8 945.4 978.0 990.8

0 0 0 0 0 0 -2 -1 -1 -1 -1 -1 -1 -1 -6 -12 -15 -18 -18 -17 -15 -15 -14 -13 -13 -12 -12 -11 -11 -10 -10 -10 -9 -9

MOMENT (kg-mm) 0 0 0 0 0 0 -391 -228 -261 -293 -326 -359 -391 -420 -2543 -5477 -7335 -9286 -9389 -9421 -8802 -9291 -9128 -8900 -9324 -8998 -9389 -8965 -9324 -8802 -9128 -9454 -8802 -8917

* Most critical fuel amount for left side most lateralC.G. condition. Weights given are nominal weights at 15 C. NOTE


412-FMS-63-5-3-2

9

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-63.2 & 63.3 & 63.4 Table 5-4.

Lateral usable fuel loading table self-sealing tanks (English) Jet B, JP-4 (6.5 Lbs/Gal)

U.S. GALLON

WEIGHT (Ibs.)

10 20 30 40 50

65 130 195 260 325 355 390 455 520 585 650 715 780 845 886 910 975 1040 1087 1105 1170 1235 1300 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2088

54.6 60 70 80 90 100 110 120 130 136.3 140 150 160 167.3 170 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 321.3

LAT CG (in.)

0 0 0 0 0 0 -0.06 -0.05 -0.05 -0.04 -0.04 -0.04 -0.03 -0.03 -0.03 -0.31 -0.50 -0.63 -0.70 -0.69 -0.65 -0.61 -0.58 -0.55 -0.53

-0.51 -0.49 -0.47 -0.45

-0.43 -0.41 -0.40 -0.39

-0.37 -0.36 -0.36

MOMENT (in-lb) 0 0 0 0 0 0

-23

-23

-26 -23 -26 -29 -23 -25 -27 -282 -488 -655 -761 -762 -761 -753 -754 -751 -758 -762 -764 -764 -761 -755 -746 -754 -761 -746 -749 -752

* Most critical fuel amount for left side most lateral C.G.condition. Weightsgiven are nominal weights at 15 C. NOTE


412-FMS-63-5-4-1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-4M. Lateral usable fuel loading table self-sealing tanks (Metric) Jet B, JP-4 (0.779 kg/I) LITERS

WEIGHT (kg)

40 80 120 160 200 206.6 240 280 320 360 400 440 480 515.6 520 560 600 633.0 640 680 720 760 800 840 880 920 960 1000 1040 1080 1120 1160 1200 1215.7

31.2 62.3 93.5 124.6 155.8 160.9 187.0 218.1 249.3 280.4 311.6 342.8 373.9 401.7 405.1 436.2 467.4 493.1 498.6 529.7 560.9 592.0 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 872.5 903.6 934.8 947.0

LAT CG (mm)

0 0 0 0 0 0 -2 -1 -1 -1 -1 -1 -1

-1 -6

-12 -15 -18 -18 -17 -15 -15 -14 -13 -13 -12 -12

-11 -11 -10 -10 -10 -9 -9

MOMENT (kg-mm)

0 0 0 0 0 0 -374 -218 -249 -280 -312 -343 -374 -402 -2430 -5235 -7011 -8876 -8974 -9005 -8413 -8881 -8725 -8507 -8912 -8600 -8974 -8569 -8912 -8413 -8725 -9036 -8413 -8523

* Most critical fuel amount for left side most lateralC.G. condition. Weights given are nominal weights at 15 C. NOTE


412-FMS-63-5-4-2

Rev.

1

11


Table 5-5. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 81.7 GAL AUX FUEL (English)

Jet A, A-1. JP-5, JP-8 (6.8 Lbs/Gal)

U.S. WEIGHT LONG CG MOMENT GALLON (lbs.) (in.) (in-lb)

10

68

139.5

20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 170 1702 180 190 200 2012 210 220 230 240 250 260 270 280 290 300 303 0 310 320 330 340 350 351 6 360 370 380 390 400 403.0

136 204 272 340 3713 408 476 544 612 680 748 816 884 952 1020 1088 1156 11574 1224 1292 1360 13682 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2060.4 2108 2176 2244 2312 2380 2390 9 2448 2516 2584 2652 2720 2740.4

139.7 139.8 139.9 139.9 1399 143.6 147.4 1498 151.8 1533 1545 155.6 1564 1572 1579 158.5 159.0 1590 1554 152.0 149.2 1489 1496 1504 151 1 151 8 152.5 153.0 153.6 154.0 154.5 154.9 155.0 154.1 1529 1517 150.6 149.5 149.4 149.8 150.3 150.8 1512 1516 151.7

9486 18999 28519 38053 47566 51942 58589 70162 81491 92902 104244 115566 126970 138258 149654 161058 172448 183804 184020 190210 196384 202912 203719 213629 224998 236320 247738 259250 270504 282010 293216 304674 315996 319362 324843 332710 340415 348187 355810 357197 366710 378155 389667 400982 412352 415719

Jet B, JP-4 (6 5 Lbs/Gal)

U.S. WEIGHT LONG CG MOMENT GALLON (lbs.) (in.) (in-lb)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 170 170.2 180 190 200 201.2 210 220 230 240 250 260 270 280 290 300 3030 310 320 330 340 350 351.6 360 370 380 390 400 403.0

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 1040 1105 11063 1170 1235 1300 13078 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 1969.5 2015 2080 2145 2210 2275 22854 2340 2405 2470 2535 2600 2619.5

139.5 139.7 139.8 1399 139.9 1399 143.6 1474 149.8 151 8 153.3 154.5 1556 156.4 157.2 157.9 158.5 159.0 159.0 1554 1520 149.2 148.9 149.6 150.4 151.1 151.8 152.5 153.0 153.6 154.0 154.5 154.9 155.0 154.1 152.9 151.7 150.6 149.5 149.4 149.8 1503 1508 151.2 151.6 151.7

9068 18161 27261 36374 45468 49651 56004 67067 77896 88803 99645 110468 121368 132158 143052 153953 164840 175695 175902 181818 187720 193960 194731 204204 215072 225895 236808 247813 258570 269568 280280 291233 302055 305273 310512 318032 325397 332826 340113 341439 350532 361472 372476 383292 394160 397378

Most critical amount for most forward C.G. condition at a weight empty below 6750 pounds has no fuel. * Most critical amount for most forward C.G. condition at a weight empty of 6750 pounds or greater. ** Most critical fuel amount for most aft C.G. condition at a weight empty below 6100 pounds. *** Most critical fuel amount for most aft C.G. condition at a weight empty of 6100 pounds or greater. Weights given are nominal weights at 15 C.

412-FMS-63-5-5-1

12

Rev.1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-5M. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 309 LITERS AUX FUEL (Metric) JetA, A-1, JP-5, JP-8 (815kg/)

LITERS

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 640 644.2 680 720 760 761.7 800 840 880 920 960 1000 1040 1080 1120 1146.7 1160 1200 1240 1280 1320 1330.6 1360 1400 1440 1480 1520 1525.2

WEIGHT LONG CG MOMENT (kg) (mm) (kg-mm) 32.6 65.2 97.8 130.4 163.0 1685 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 5216 525 0 554.2 586.8 619.4 620.8 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 912.8 934.6 945.4 978.0 1010.6 1043.2 1075.8 1084.4 1108.4 11410 11736 1206.2 1238.8 1243.0

3543 3548 3551 3553 3553 3553 3691 3772 3830 3876 3912 3942 3967 3988 4008 4023 4036 4038 3950 3861 3785 3782 3800 3823 3843 3861 3879 3875 3906 3919 3931 3938 3926 3879 3872 3830 3801 3794 3805 3818 3830 3840 3851 3852

115502 231330 347288 463311 579139 598540 721960 860770 998864 1137218 1275312 1413601 1551890 1690114 1829251 1967247 2105178 2120043 2189090 2265635 2344429 2347811 2477600 2617226 2756200 2894978 3034930 3158125 3310726 3449504 3588217 3680299 3711640 3793662 3913043 3995456 4089116 4114362 4217462 4356338 4494888 4631808 4770619 4788182

Most critical amount for most forward C.G. condition * Most critical amount for most forward C.G. condition ** Most critical fuel amount for most aft C.G. condition ***Most critical fuel amount for most aft C.G. condition Weights given are nominal weights at 15 C.

Jet B, JP-4 (.779 kg/1)

LITERS

40 80 120 160 200 2067 240 280 320 360 400 440 480 520 560 600 640 644.2 680 720 760 761.7 800 840 880 920 960 1000 1040 1080 1120 1146.7 1160 1200 1240 1280 1320 1330.6 1360 1400 1440 1480 1520 1525.2

WEIGHT LONG CG MOMENT (kg) (mm) (kg-mm) 31.2 62.3 93.5 124.6 155.8 161.0 1870 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 4674 498.6 501.8 529.7 560.9 592.0 593.4 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 872.5 893.3 903.6 934.8 9660 997 1 1028.3 1036.5 1059.4 1090.6 1121.8 1152.9 1184.1 1188.1

3543 3548 3551 3553 3553 3553 3691 3772 3830 3876 3912 3942 3967 3988 4008 4023 4036 4038 3950 3861 3785 3782 3800 3823 3843 3861 3879 3875 3906 3919 3931 3938 3926 3879 3872 3830 3801 3794 3805 3818 3830 3840 3851 3852

110400 221111 331947 442846 553557 572102 690069 822749 954742 1086985 1218979 1351160 1483341 1615459 1748450 1880350 2012188 2026397 2092394 2165558 2240871 2244104 2368160 2501618 2634453 2767101 2900871 3018625 3164485 3297133 3429719 3517734 3547691 3626089 3740197 3818970 3908492 3932623 4031169 4163911 4296341 4427213 4559892 4576680

at a weight empty below 3062 kilograms has no f at a weight empty of 3062 kilograms or greater. at a weight empty below 2767 kilograms. at a weight empty of 2767 kilograms or greater.

412-FMS-63-5-5-2

Rev.1

13


Table 5-6. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 81.7 GAL AUX FUEL (LH) (ENGLISH) Jet A, A-1, JP-5, JP-8 (6 8 Lbs/Gal)

10 20 30 40 50 546 60 70 80 90 100 110 120 130 140 150 160 170 1702 180 190 200 201.2 210 220 230 240 250 260 270 280 290 300 303.0 310 320 330 340 350 3516 360 370 380 390 400 403 0

68 136 204 272 340 371.3 408 476 544 612 680 748 816 884 952 1020 1088 1156 1157.4 1224 1292 1360 1368.2 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2060.4 2108 2176 2244 2312 2380 2390.9 2448 2516 2584 2652 2720 2740.4

0 0 0 0 0 0 -0.13 -0.72 -1.78 -2.63 -3.30 -3.84 -4.29 -4.67 -5.00 -5.27 -5.50 -5.71 -5.71 -5 72 -5.56 -5.41 -5.39 -5.54 -5 69 -584 -5 97 -608 -6 19 -629 -639 -648 -6 56 -6.58 -6.43 -6.23 -6.04 -5 86 -5 70 -5.67 -5 75 -5.84 -5 92 -600 -6.07 -6.09

0 0 0 0 0 0 -53 -343 -968 -1610 -2244 -2872 -3501 -4128 -4760 -5375 -5984 -6601 -6609 -7001 -7184 -7358 -7374 -7911 -8512 -9134 -9743 -10336 -10944 -11548 -12167 -12779 -13382 -13557 -13554 -13556 -13554 -13548 -13566 -13556 -14076 -14693 -15297 -15912 -16510 -16689

Jet B. JP-4 (6.5 Lbs/Gal)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 170 170.2 180 190 200 201.2 210 220 230 240 250 260

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 1040 1105 1106.3 1170 1235 1300 1307.8 1365 1430 1495 1560 1625 1690

0 0 0 0 0 0 -0.13 -072 -1.78 -2 63 -3.30 -3.84 -429 -4.67 -5.00 -5.27 -5.50 -5.71 -5.71 -5 72 -5.56 -541 -5.39 -5.54 -5.69 -5.84 -5 97 -6 08 -6 19

* Most critical fuel amount for left side most lateral C.G. condition. Weights given are nominal weights at 15 C.

412-FMS-63-5-6-1

14

Rev.1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-6M. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 309 LITERS AUX FUEL (LH) (Metric) Jet A, A-1 JP-5, JP-8 (815kg/1)

LITERS

40 80 120 160 200 2067 240 280 320 360 400 440 480 520 560 600 640 644.2 680 720 760 761.7 800 840 880 920 960 1000 1040 1080 1120 1146.7 1160 1200 1240 1280 1320 13306 1360 1400 1440 1480 1520 1525.2

WEIGHT LAT CG (kg) (mm)

32.6 65.2 97.8 130.4 163.0 168.5 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 525.0 554 2 586 8 619.4 620.8 652.0 684.6 717.2 749.8 782.4 815.0 8476 880 2 912.8 934.6 945.4 978.0 1010.6 1043.2 10758 10844 1108.4 1141.0 1173.6 12062 12388 1243.0

0 0 0 0 0 0 -6 -28 -56 -76 -92 -105 -116 -126 -132 -139 -145 -145 -145 -141 -137 -137 -141 -145 -149 -152 -156 -158 -161 -164 -166 -167 -165 -160 -155 -150 -145 -144 -146 -148 -150 -152 -154 -155

Jet B, JP-4 ( 779 kg/I)

MOMENT (kg-mm)

LITERS

0 0 0 0 0 0 -1174 -6390 -14605 -22298 -29992 -37653 45379 -53399 -60245 -67971 -75632 -76128 -80359 -82739 -84858 -85048 -91932 -99267 -106863 -113970 -122054 -128770 -136464 -144353 -151525 -156072 -155991 -156480 -156643 -156480 -155991 -156159 -161826 -168868 -176040 -183342 -190775 -192671

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 640 644.2 680 720 760 761.7 800 840 880 920 960 1000 1040 1080 1120 1146.7 1160 1200 1240 1280 1320 1330.6 1360 1400 1440 1480 1520 15252


WEIGHT LATCG (mm) (kg)

31.2 62.3 93.5 124.6 155.8 161.0 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 498.6 501.8 529 7 560 9 592.0 593.4 623.2 654.4 685.5 716.7 747.8 779.0 810.2 841.3 872.5 893.3 903.6 934.8 966.0 997.1 1028.3 1036.5 10594 1090.6 1121.8 1152.9 1184.1 1188.1

0 0 0 0 0 0 -5 -25 -55 -75 -91 -105 -116 -126 -133 -140 -147 -145 -145 -141 -137 -137 -142 -146 -150 -152 -155 -158 -161 -164 -166 -167 -165 -158 -156 -150 -145 -144 -146 -149 -151 -153 -154 -155

MOMENT (kg-mm)

0 0 0 0 0 0 -935 -5453 -13710 -21033 -28356 -35990 43375 -51040 -58020 -65436 -73288 -72766 -76809 -79084 -81109 -81291 -88494 -95537 -102828 -108935 -115915 -123082 -130436 -137976 -144832 -149178 -149101 -147698 -150690 -149568 -149101 -149261 -154678 -162499 -169386 -176397 -182348 -184160

412-FMS-63-5-6-2

Rev. 1

15


Table 5-7. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 81.7 GAL AUX FUEL (RH) (Engllsh) Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal)

568 5.79 5.82 5.69 5.51 534 5.19 5.04 5.01 5.11 5.21 531 5.40 5.48 551

11201 11812 11992 11995 11990 11983 11999 11995 11978 12509 13108 13721 14321 14906 15100

Jet B. JP-4 (6 5 Lbs/Gal)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 170 170.2 180 190 200 201.2 210 220 230 240 250 260 270 280 290 300 303.0 310 320 330 340 350 351.6 360 370 380 390 400 4030

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 1040 1105 11063 1170 1235 1300 1307.8 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 1969.5 2015 2080 2145 2210 2275 22854 2340 2405 2470 2535 2600 2619.5

0 0 0 0 0 0 0.01 0.62 168 2 55 3.23 3.77 4.23 4.61 4.95 5.22 5.46 5.66 5.66 5.03 4.63 4.42 423 442 463 483 499 5.14 5.30 5.43 556 568 5.79 5.82 5.69 5.51 534 519 5.04 5.01 5.11 5.21 5.31 5 40 5.48 5.51

0 0 0 0 0 0 4 282 874 1492 2100 2696 3299 3895 4505 5090 5678 6254 6262 5885 5718 5746 5532 6033 6621 7221 7784 8353 8957 9530 10119 10707 11291 11462 11465 11461 11454 11470 11466 11450 11957 12530 13116 13689 14248 14433

Most critical fuel amount for right side most lateral C.G. condition. Weights given are nominal weights at 15 C.

412-FMS-63-5-7-1

16

Rev.

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-7M. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 309 LITERS AUX FUEL (RH) (Metric)

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 640 644.2 680 720 760 761.7 800 840 880 920 960 1000 1040 1080 1120 1146 7 1160 1200 1240 1280 1320 1330.6 1360 1400 1440 1480 1520 15252

326 652 97.8 130.4 163.0 168 5 1956 228 2 260 8 293 4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 525 0 554 2 586 8 6194 620.8 652.0 684.6 717.2 749.8 782.4 815.0 847.6 880.2 912.8 934.6 945.4 978.0 1010.6 1043.2 1075.8 1084.4 1108.4 1141.0 1173.6 12062 1238.8 1243.0

0 0 0 0 0 0 3 25 53 74 90 103 114 123 131 138 143 144 128 117 108 107 113 119 124 128 132 136 140 143 146 148 146 139 141 137 128 127 129 132 135 137 140 140

0 0 0 0 0 0 497 5705 13822 21712 29340 36936 44597 52127 59788 67482 74589 75603 70938 68656 66895 66424 73676 81467 88933 95974 103277 110840 118664 125869 133269 138315 138028 135942 142495 142918 137702 137724 142984 150612 158436 165249 173432 174025

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 640 644.2 680 720 760 761.7 800 840 880 920 960 1000 1040 1080 1120 1146.7 1160 1200 1240 1280 1320 1330.6 1360 1400 1440 1480 1520 1525.2

312 623 93 5 1246 1558 161.0 1870 218 1 249 3 280 4 311 6 342.8 373.9 405.1 436.2 467.4 498 6 501.8 529 7 560.9 592 0 5934 623 2 6544 685 5 716.7 747.8 779.0 8102 841 3 872.5 893.3 9036 934.8 9660 9971 10283 10365 10594 10906 11218 11529 1184.1 1188.1

0 0 0 0 0 0 3 25 53 74 90 103 114 123 131 138 143 144 128 117 108 107 113 119 124 128 132 136 140 143 146 148 146 139 141 137 128 127 129 132 135 137 140 140

0 0 0 0 0 0 475 5453 13212 20753 28044 35304 42627 49825 57147 64501 71294 72264 67804 65623 63940 63490 70422 77869 85004 91735 98715 105944 113422 120309 127382 132205 131931 129937 136200 136605 131620 131640 136668 143959 151438 157950 165771 166338

* Most critical fuel amount for right side most lateral C.G. condition. Weights given are nominal weights at 15 C.

412-FMS-63-7-2

Rev.1

17


Table 5-8. USABLE FUEL LOADING TABLE WITH SELF-SEALING (ENGLISH)

TANKS AND 163.4 GAL AUX FUEL

Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal)

U.S.

WEIGHT

LONG CG

MOMENT

U.S.

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 204.1 210 220 230 235 2 240 250 260 270 280 290 300 310 320 330 340 350 360 3686 370 380 390 400 410 4172 420 430 440 450 460 470 480 4846

Most critical no fuel.

* Most critical ** Most critical

amount for most forward

C.G. condition

amount for most forward C.G. condition fuel amount for most aft C.G. condition.

WEIGHT

65 130 195 260 325 3549 390 455 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1326 7 1365 1430 1495 1528.8 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2395.9 2405 2470 2535 2600 2665 2711.8 2730 2795 2860 2925 2990 3055 3120 3149.9

LONGCG

MOMENT

139.5 139.7 1398 139.9 139.9 1399 143.6 147.3 149.2 150.7 151 8 152.8 153.6 1543 154.8 1554 155.8 156.2 156.6 1570 157.2 1573 1555 152.7 1501 148 9 149.2 149 7 150.2 1506 151.1 151 5 151.9 152.3 152.6 152.9 153.2 1535 153.7 153.9 153.8 152.8 151.8 1508 149.9 149.3 149.4 149.6 149.9 150.2 150.5 150.8 151.0 151.2

9068 18161 27261 36374 45468 49651 56004 67022 77584 88160 98670 109252 119808 130384 140868 151515 162032 172601 183222 193895 204360 208682 212258 218361 224400 227638 232752 243263 253838 264303 275002 285578 296205 306885 317408 327971 338572 349213 359658 368729 369889 377416 384813 392080 399484 404872 407862 418132 428714 439335 449995 460694 471120 476265

at a weight empty below 6920 pounds has at a weight empty of 6920 pounds

or greater.

Weights given are nominal weights at 15 C. 412-FMS-63-5-8-1

Rev. 1

BHT-412-FMS-63.2

& 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-8M. USABLE FUEL LOADINGTABLE WITH SELF-SEALINGTANKS AND 618 LITERSAUX FUEL (Metric)

Jet A, A-1, JP-5, JP-8 (.815kgfl)

LITERS

WEIGHT LONG CG (mm) (kg)

Jet B, JP-4 (.779 kg/1)

MOMENT (kg-mm)

40 80 120 160 200

32.6 652 97.8 130.4 1630

3543 3548 3551 3553 3553

115502 231330 347288 463311 579139

206.7

1685

3553

598540

195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 4564 489.0 521.6 554.2 586.8 619.4 629.8 652.0 684.6 717.2 725.5 7498 782 4 8150 847.6 880 2 912.8 945.4 978.0 10106 1043 2 1075.8 11084 1137.1 1141.0 1173.6 12062 1238.8 1271.4 1287.0 1304.0 1336.6 1369.2 1401.8 1434.4 1467.0 1495.0

3691 3764 3810 3843 3866 3894 3917 3929 3945 3957 3967 3978 3988 3993 3996 3952 3866 3797 3782 3792 3807 3820 3810 3846 3856 3864 3874 3882 3890 3897 3904 3910 3906 3879 3852 3827 3803 3792 3795 3802 3813 3820 3825 3833 3840

240 280 320 360 400 440 480 520 560 600 640 680 720 760 772.7 800 840 880 890.2 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1395.2 1400 1440 1480 1520 1560 15792 1600 1640 1680 1720 1760 1800 1834.4

721960 858945 993648 1127536 1260316 1396388 1532330 1665110 1800498 1934973 2069187 2204608 2340158 2473264 2516483 2576704 2646664 2723208 2743890 2843242 2978597 3113300 3229356 3385249 3519757 3653026 3788772 3923149 4058048 4192393 4327194 4446014 4456746 4552394 4646282 4740888 4835134 4880486 4948680 5081753 5220760 5354876 5486580 5623011 5740938

WEIGHT LONG CG (kg) (mm) LITERS 40 80 120 160 200

2067

240 280 320 360 400 440 480 520 560 600 640 680 720 760 772.7 800 840 880 890.2 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1395.2 1400 1440 1480 1520 1560 1579.2 1600 1640 1680 1720 1760 1800 1834.4

MOMENT (kg-mm)

31.2 62.3 93.5 124.6 1558

3543 3548 3551 3553 3553

110400 221111 331947 442846 553557

161.0

3553

572102

187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 4362 467.4 498.6 529.7 560.9 592.0 601.9 623.2 654.4 685.5 693.5 716.7 747.8 779.0 8102 841.3 872.5 903.6 934.8 966 0 9971 1028.3 1059.4 1086.9 1090.6 1121.8 1152.9 1184.1 1215.2 1230.2 12464 12776 1308.7 1339.9 1371.0 1402.2 1429.0

3691 3764 3810 3843 3866 3894 3917 3929 3945 3957 3967 3978 3988 3993 3996 3952 3866 3797 3782 3792 3807 3820 3810 3846 3856 3864 3874 3882 3890 3897 3904 3910 3906 3879 3852 3827 3803 3792 3795 3802 3813 3820 3825 3833 3840

690069 821004 949757 1077731 1204646 1334707 1464645 1591559 1720967 1849502 1977788 2107226 2236789 2364016 2405325 2462886 2529756 2602919 2622688 2717651 2847027 2975780 3086710 3235717 3364283 3491665 3621415 3749857 3878797 4007207 4136054 4249626 4259884 4351307 4441048 4531474 4621558 4664906 4730088 4857283 4990149 5118342 5244228 5374633 5487351

Most critical amount for most forward CG. condition at a weight empty below 3139 kilogramshas n Most critical amount for most forward C.G. condition at a weight empty of 3062 kilograms or greater ** Most critical fuel amount for most aft C.G. condition Weights given are nominal weights at 15 C.

412-fms-63-5-8-2

Rev.1

19


Table 5-9. USABLE FUEL LOADING TABLE WITH SELF-SEALING (English) Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal) U.S. WEIGHT GALLONS (lbs)

10 20 30 40 50 546 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 204.1 210 220 230 235.2 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 390 400 410 420 430 440 450 460 470 480 484.6

68 136 204 272 340 371.3 408 476 544 612 680 748 816 884 952 1020 1088 1156 1224 1292 1360 1387.9 1428 1496 1564 1599.4 1632 1700 1768 1836 1904 1972 2040 2108 2176 2244 2312 2380 2448 2516 2584 2652 2720 2788 2856 2924 2992 3060 3128 3196 3264 3295.3

LAT CG (in.)

0 0 0 0 0 0 -0.08 -0.04 -0.04 -0.04 -0.04 -0.03 -0.03 -003 -0.03 -0 02 -0.02 -002 -0 02 -0.02 -0.02 -0.02 -0.25 -035 -0.45 -0.50 -0.48 -0.46 -044 -0.43 -0.42 -0.40 -0.39 -0.37 -0.37 -0.36 -0.34 -0.33 -0.32 -032 -031 -0.30 -0.30 -0.29 -0.28 -0.27 -0.26 -0.26 -025 -0 25 -0 24 -0.24

MOMENT (in-lb)

0 0 0 0 0 0 -33 -19 -22 -24 -27 -22 -24 -27 -29 -20 -22 -23 -24 -26 -27 -28 -357 -524 -704 -800 -783 -782 -778 -789 -800 -789 -796 -780 -805 -808 -786 -785 -783 -805 -801 -796 -816 -809 -800 -789 -778 -796 -782 -799 -783 -791

TANKS AND 163.4 GAL AUX FUEL Jet B. JP-4 (6.5 Lbs/Gal)

U.S. WEIGHT GALLONS (lbs.)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 204.1 210 220 230 235.2 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 390 400 410 420 430 440 450 460 470 480 484.6

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1326.7 1365 1430 1495 1528.8 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2405 2470 2535 2600 2665 2730 2795 2860 2925 2990 3055 3120 3149.9

LAT CG (in.)

0 0 0 0 0 0 -0.08 -004 -004 -0.04 -0.04 -0.03 -0 03 -0.03 -0.03 -0.02 -002 -0.02 -0.02 -002 -0.02 -0.02 -0.25 -0.35 -0.45 -0.50 -0.48 -0.46 -0.44 -043 -0.42 -0.40 -0.39 -0.37 -0.37 -0.36 -0.34 -0.33 -0.32 -0.32 -0.31 -0.30 -0.30 -0.29 -0.28 -0.27 -0.26 -0.26 -0.25 -0.25 -0.24 -0.24

MOMENT (in-lb)

0 0 0 0 0 0 -31 -18 -21 -23 -26 -21 -23 -25 -27 -20 -21 -22 -23 -25 -26 -27 -341 -501 -673 -764 -749 -748 -744 -755 -764 -754 -761 -746 -770 -772 -751 -751 -749 -770 -766 -761 -780 -773 -764 -755 -744 -761 -748 -764 -749 -756


412-FMS-63-5-9-1

20

Rev. 1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-9M. USABLE FUEL LOADINGTABLEWITH SELF-SEALINGTANKSAND 618 LITERSAUX FUEL (Metric)

Jet A, A-1, JP-5, JP-8 (815kg/1)

LITERS

40 80 120 160 200 206.7 240 280 320 360 400

(kg)

326 652 978 1304 163.0 168.5 1956 228.2 260.8 293.4 326.0

440

358.6

480 520 560 600 640 680 720 760 772.7 800 840 880 890.2 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1395.2 1400 1440 1480 1520 1560 1579.2 1600 1640 1680 1720 1760 1800 1834.4

391 2 423 8 456 4 489.0 521.6 554.2 586.8 619.4 629 8 652.0 684.6 717.2 725.5 749 8 782.4 8150 847.6 880.2 912.8 9454 9780 1010.6 1043.2 1075 8 11084 1137.1 11410 1173.6 1206.2 1238 8 12714 1287.0 13040 13366 1369.2 14018 1434.4 1467.0 1495.0

(kg-mm)

(mm)

0 0 0 0 0 0 -1 -1 -1 -1 -1

0 0 0 0 0 0 -196 -228 -261 -293 -326

-1

-359

-1 -1 -1 -1 -1 -1 -1 -1 -1 -7 -9 -12 -13 -12 -12 -11 -11 -10 -10 -10 -9 -9 -9 -8 -8 -8 -8 -8 -8 -7 -7 -7 -7 -7 -7 -6 -6 -6 -6

-391 -424 -456 -489 -522 -554 -587 -619 -630 -4564 -6161 -8606 -9432 -8998 -9389 -8965 -9324 -8802 -9128 -9454 -8802 -9095 -9389 -8606 -8867 -9097 -9128 -9389 -9650 -8672 -8900 -9009 -9128 -9356 -9584 -8411 -8606 -8802 -8970

LITERS

(kg)

(kg-mm)

(mm)

0 0 0 0 0 0 -187 -218 -249 -280 -312

31.2 62.3 93.5 1246 155.8 161 0 187.0 2181 249.3 280.4 311.6

0 0 0 0 0 0 -1 -1 -1 -1 -1

440

342.8

-1

-343

480 520 560 600 640 680 720 760 772.7 800 840 880 890.2 920 960 1000 1040 1080 1120 1160 1200 1240 1280 1320 1360 1395.2 1400 1440 1480 1520 1560 15792 1600 1640 1680 1720 1760 1800 1834.4

373.9 405.1 436.2 467 4 498.6 529.7 560.9 592 0 601 9 623 2 654.4 685 5 6935 716.7 747 8 7790 810.2 841.3 872.5 9036 9348 9660 997 1 1028 3 1059.4 10869 1090.6 1121 8 1152.9 1184.1 12152 12302 12464 1277 6 1308 7 1339.9 1371.0 1402.2 1429.0

-1 -1 -1 -1 -1 -1 -1 -1 -1 -7 -9 -12 -13 -12 -12 -11 -11 -10 -10 -10 -9 -9 -9 -8 -8 -8 -8 -8 -8 -7 -7 -7 -7 -7 -7 -6 -6 -6 -6

-374 -405 -436 -467 -499 -530 -561 -592 -602 -4362 -5889 -8226 -9015 -8600 -8974 -8569 -8912 -8413 -8725 -9036 -8413 -8694 -8974 -8226 -8476 -8695 -8725 -8974 -9223 -8289 -8507 -8611 -8725 -8943 -9161 -6039 -8226 -8413 -8574

40 80 120 160 200 206.7 240 280 320 360 400

*Weights given are nominalweightsat 15 C. Weights given are nominalweightsat 15 C.

Rev. 1

21


Table 5-10. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 16.3 GAL AUX FUEL (English) Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal) Jet B, JP-4 (6 5 Lbs/Gal)

U.S. WEIGHT LONG CG MOMENT GALLON (lbs.) (In.) (In-lb)

10 20 30 40 50 546 60 70 80 90 100 110 120 130 140 150 152.6 160 170 180 183.6 190 200 210 220 230 240 250 253.6 260 270 280 290 300 302.2 310 320 330 3376

68 136 204 272 340 3713 406 476 544 612 680 748 816 884 952 1020 1037.7 1088 1156 1224 12485 1292 1360 1428 1496 1564 1632 1700 1724.5 1768 1836 1904 1972 2040 2055.0 2108 2176 2244 2295.7

139.5 139.7 139.8 139.9 139.9 139.9 143.2 1474 150.5 152.7 1545 155.9 157.1 158.2 159.1 1599 1602 157.1 153.3 1500 1489 149.8 151.1 1523 153.3 1543 1552 156.0 1563 155.2 1537 152.3 1510 149.8 149.5 150.2 150.9 151 7 152.2

9486 18999 28519 38053 47566 51942 58426 70162 81872 93452 105060 116613 128194 139849 151463 163098 166236 170925 177215 183600 185899 193542 205496 217484 229337 241325 253286 265200 269536 274394 282193 289979 297772 305592 307217 316622 328358 340415 349402

U.S. WEIGHT LONG CG MOMENT GALLON (Ibs.) (in.) (in-lb)

*

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 1526 160 170 180 183.6 190 200 210 220 230 240 250 253.6 260 270 280 290 300 302.2 310 320 330 337.6

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 991.9 1040 1105 1170 1193.4 1235 1300 1365 1430 1495 1560 1625 1648.4 1690 1755 1820 1885 1950 1964.3 2015 2080 2145 2194.4

1395 1397 139.8 139.9 1399 139.9 1432 1474 150.5 1527 1545 155.9 157.1 158.2 159.1 159.9 160.2 1571 153.3 150.0 148.9 149 8 151.1 152.3 1533 154.3 155.2 156.0 156.3 1552 153.7 152 3 151.0 149.8 149.5 150.2 150.9 151.7 152.2

9068 18161 27261 36374 45468 49651 55848 67067 78260 89330 100425 111469 122538 133679 144781 155903 158902 163384 169397 175500 177697 185003 196430 207890 219219 230679 242112 253500 257645 262288 269744 277186 284635 292110 293663 302653 313872 325397 333988

Most critical amount for most forward C.G. condition at a weight empty below 6580 pounds has no fuel. Most critical amount for most forward C.G. condition at a weight empty of 6580 pounds or greater. Most critical fuel amount for most aft C.G. condition at a weight empty below 6650 pounds. Most critical fuel amount for most aft C.G. condition at a weight empty of 6650 pounds or greater. Weights given are nominal weights at 15 C.

412-FMS-63-5-10-1

22

Rev. 1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table5-10M. USABLEFUELLOADINGTABLEWITHSELF-SEALINGTANKSAND 62 LITERSAUX FUEL (Metric) Jet B, JP-4 ( 779 kg/1) Jet A, A-1, JP-5,JP-8 (.815kg/1)

LITERS

WEIGHTLONGCG MOMENT (kg-mm) (mm) (kg)

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 577.4 600 640 680 694.9 720 760 800 840 880 920 959.8 960 1000 1040 1080 1120 11438 1160 1200 1240 1277.7

32.6 652 97.8 130.4 1630 168.5 195.6 228.2 260.8 293.4 326.0 358.6 3912 423 8 456.4 470.6 489.0 521 6 554.2 566.3 586.8 619.4 652.0 684.6 717.2 749.8 782 2 782.4 8150 847.6 880.2 912.8 932.2 945.4 978.0 1010.6 1041.3

3543 3548 3551 3553 3553 3553 3679 3777 3850 3904 3946 3981 4010 4034 4057 4069 4005 3903 3813 3782 3806 3840 3871 3900 3925 3948 3970 3969 3927 3887 3850 3816 3797 3806 3828 3846 3867

115502 231330 347288 463311 579139 598540 719612 861911 1004080 1145434 1286396 1427587 1568712 1709609 1851615 1914794 1958445 2035805 2113165 2141911 2233361 2378496 2523892 2669940 2815010 2960210 3105481 3105346 3200505 3294621 3388770 3483245 3539552 3598192 3743784 3886768 4026806

LITERS

WEIGHTLONGCG MOMENT (kg-mm) (mm) (kg)

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 ** 577.4 600 640 680 694.9 720 760 800 840 880 920 959 8 960 1000 1040 1080 1120 1143.8 1160 1200 1240 1277.7

31.2 62.3 93.5 124.6 155.8 161.0 187.0 218.1 249.3 2804 311.6 342.8 373.9 405.1 436.2 449.8 467 4 498 6 529.7 541.3 560.9 592.0 623.2 654.4 685.5 716.7 747.7 747.8 779 0 810.2 841.3 872.5 891.0 903.6 934.8 966.0 995.3

3543 3548 3551 3553 3553 3553 3679 3777 3850 3904 3946 3981 4010 4034 4057 4069 4005 3903 3813 3782 3806 3840 3871 3900 3925 3948 3970 3969 3927 3887 3850 3816 3797 3806 3828 3846 3867

110400 221111 331947 442846 553557 572102 687826 823839 959728 1094838 1229574 1364528 1499419 1634093 1769826 1830214 1871937 1945880 2019822 2047299 2134709 2273434 2412407 2552004 2690666 2829453 2968306 2968177 3059133 3149092 3239082 3329384 3383204 3439254 3578414 3715082 3848935

Most critical amount for most forward C.G.condition at a weightempty below2984 kilograms has no fuel. Mostcriticalamountfor most forwardC.G.conditionat a weightempty of 2984 kilogramsor greater. Mostcriticalfuel amountfor most aft C.G. conditionat a weightempty below3016 kilograms. Mostcriticalfuel amountfor most aft C.G.conditionat a weightempty of 3016 kilogramsor greater. Weightsgivenare nominalweightsat 15 C.

412-FMS-63-5-10-2

Rev. 1

23


Table 5-11.USABLE FUEL LOADINGTABLEWITH SELF-SEALINGTANKSAND 16.3 GAL AUX FUEL(LH) (English)

Jet A, A-1, JP-5,JP-8 (6.8 Lbs/Gal)


U.S. WEIGHT LATCG MOMENT GALLON (lbs.) (in.) (In-lb)

U.S. WEIGHT LATCG MOMENT GALLON (lbs.) (in.) (In-lb)

10 68 20 136 30 204 40 272 50 340 54.6 371.3 60 408 70 476 80 544 90 612 100 680 110 748 120 816 130 884 140 952 150 1020 152.6 1037.7 160 1088 170 1156 180 1224 1836 12485 190 1292 200 1360 210 1428 220 1496 230 1564 240 1632 250 1700 253.6 1724.5 260 1768 270 1836 280 1904 290 1972 300 2040 302.2 2055.0 310 2108 320 2176 330 2244 3376 22957

0 0 0 0 0 0 -0.17 -0 69 -1 02 -1 42 -1.75 -2 00 -2.22 -2.40 -2.55 -2.58 -2.54 -2 74 -2.74 -2.73 -2.72 -2.63 -2.49 -2.37 -2.27 -2.20 -2.08 -2.00 -1.97 -1.92 -1.85 -1.78 -1.72 -1.67 -1.65 -1.61 -1.57 -1.52 -1.48

0 0 0 0 0 0

-69 -328 -555 -869 -1190 -1496 -1812 -2122 -2428 -2632 -2636 -2981 -3167 -3342 -3396 -3398 -3386 -3384 -3396 -3441 -3395 -3400 -3397 -3395 -3397 -3389 -3392 -3407 -3391 -3394 -3416 -3411 -3398

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 1526 160 170 180 1836 190 200 210 220 230 240 250 253.6 260 270 280 290 300 3022 310 320 330 3376

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 991 9 1040 1105 1170 1193.4 1235 1300 1365 1430 1495 1560 1625 1648.4 1690 1755 1820 1885 1950 1964.3 2015 2080 2145 21944

0 0 0 0 0 0 -0.17 -0.69 -1.02 -1.42 -1.75 -200 -2 22 -2.40 -2.55 -2.58 -2.54 -2.74 -2.74 -2.73 -2.72 -2.63 -2.49 -2 37 -2.27 -2.20 -2.08 -2.00 -1.97 -1.92 -1.85 -1.78 -1.72 -1.67 -1.65 -1.61 -1.57 -1.52 -1.48

0 0 0 0 0 0

-66 -314 -530 -831 -1138 -1430 -1732 -2028 -2321 -2516 -2519 -2850 -3028 -3194 -3246 -3248 -3237 -3235 -3246 -3289 -3245 -3250 -3247 -3245 -3247 -3240 -3242 -3257 -3241 -3244 -3266 -3260 -3248

* Most critical fuel amount for left side most lateral C.G.condition. Weights given are nominal weights at 15 C.

412-FMS-63-5-11-1

24

Rev. 1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-11M. USABLE FUELLOADINGTABLE WITHSELF-SEALINGTANKSAND 62 LITERSAUX FUEL(LH) (Metric)

Jet B, JP-4 (.779 kg/1)

JetA, A-1,JP-5, JP-8(815kg/1)

LITERS

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 577.4 600 640 680 694.9 720 760 800 840 880 920 959.8 960 1000 1040 1080 1120 1143.8 1160 1200 1240 1277.7

WEIGHT LATCG MOMENT (mm) (kg-mm) (kg)

32.6 65.2 97.8 130.4 163.0 168.5 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 470.6 489.0 521.6 554.2 566.3 586.8 619.4 652.0 684.6 717.2 749.8 782.2 782.4 815.0 847.6 880.2 912.8 932.2 945.4 978.0 1010.6 1041.3

0 0 0 0 0 0

-8 -22 -31 -41 48 -54 -60 -64 -66 -64 -70 -70 -69 -69 -67 -63 -60 -57 -54 -52 -50 -50 -48 -46 -44 43 -42 41 -40 -39 -38

0 0 0 0 0 0

-1565 -5020 -8085 -12029 -15648 -19364 -23472 -27123 -30122 -30117 -34230 -36512 -38240 -39078 -39316 -39022 -39120 -39022 -38729 -38990 -39112 -39120 -39120 -38990 -38729 -39250 -39152 -38761 -39120 -39413 -39570

LITERS

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 577.4 600 640 680 694.9 720 760 800 840 880 920 959.8 960 1000 1040 1080 1120 1143.8 1160 1200 1240 1277.7

WEIGHT LAT CG MOMENT (kg) (mm) (kg-mm)

31.2 62.3 93.5 124.6 155.8 161.0 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436 2 449.8 467.4 498.6 529.7 541.3 560.9 592.0 623.2 654.4 685.5 716.7 747.7 747 8 779.0 810.2 841.3 872.5 891.0 903.6 934.8 966.0 995.3

0 0 0 0 0 0

0 0 0 0 0 0

-8 -22 -31 -41 -48 -54 -60 -64 -66 -64 -70 -70 -69 -69 -67 -63 -60 -57 -54 -52 -50 -50 -48 46 -44 -43 -42 -41 -40 -39 -38

-1496 4799 -7728 -11498 -14957 -18509 -22435 -25925 -28792 -28787 -32718 -34899 -36551 -37352 -37579 -37299 -37392 -37299 -37018 -37267 -37384 -37392 -37392 -37267 -37018 -37517 -37423 -37049 -37392 -37672 -37822


412-FMS-63-5-11-2

Rev. 1

25

T N I R P N O R I O T F C U 0 T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-63.2 & 63.3 & 63.4

Table 5-12. USABLE FUELLOADINGTABLE WITH SELF-SEALINGTANKS AND 16.3 GALAUX FUEL(RH) (English)

U.S. WEIGHT LAT CG MOMENT GALLON (Ibs.) (in.) (in-lb) 10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 152.6 160 170 180 183.6 190 200 210 220 230 240 250 253.6 260 270 280 290 300 302.2 310 320 330 337 6

68 136 204 272 340 371 3 408 476 544 612 680 748 816 884 952 1020 1037 7 1088 1156 1224 12485 1292 1360 1428 1496 1564 1632 1700 17245 1768 1836 1904 1972 2040 2055.0 2108 2176 2244 2295.7

0 0 0 0 0 0

0.06 0.60 0.93 1.34 1.67 1.93 216 2.35 2 50 2.53 249 2.05 1.77 1.53 1.45 1.40 1.33 127 1.21 1.16 1.11

1.07 1.05 102

0.99

095 0.92 0.89 0.88 0.86 0.83 0.80 0.79

0 0 0 0 0 0 24 286 506 820

1136 1444 1763 2077 2380 2581 2584 2230 2046 1873 1810 1809 1809 1814 1810 1814 1812 1819 1811 1803 1818 1809 1814 1816 1808 1813 1806 1795 1814

U.S. WEIGHT LATCG MOMENT (In.) (In-lb) GALLON (lbs.) 10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 1526 160 170 180 183.6 190 200 210 220 230 240 250 253.6 260 270 280 290 300 302.2 310 320 330 337.6

65 130 195 260 325 3549 390 455 520 585 650 715 780 845 910 975 9919 1040 1105 1170 11934 1235 1300 1365 1430 1495 1560 1625 1648.4 1690 1755 1820 1885 1950 1964.3 2015 2080 2145 2194.4

0 0 0 0 0 0 0.06 0.60 0.93 1.34 1.67 1.93 2.16 2 35 2 50 2.53 2.49 2.05 1.77 1.53 1.45 1.40 33 1.27 1.21 1.16 1.11

1.07 1.05

1.02 0.99 0.95

0.92 0.89 0.88 0.86 0.83 080 0.79

0

0 0 0 0 0

23 273 484 784 1086 1380 1685 1986 2275 2467 2470 2132 1956 1790 1730 1729 1729 1734 1730 1734 1732 1739 1731 1724 1737 1729 1734 1736 1729 1733 1726 1716 1734

* Most critical fuel amount for right side most lateralC.G. condition. Weights given are nominal weights at 15 C.

412-FMS-63-5-12-1

26

Rev. 1


Table 5-12M. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 62 LITERS AUX FUEL (RH) (Metric) Jet A, A-1, JP-5, JP-8 (.815kg/1) Jet B, JP-4 (.779 kg/1)

LITERS

WEIGHT (kg)

LAT CG (mm)

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 577.4 600 640 680 6949 720 760 800 840 880 920 959.8 960 1000 1040 1080 1120 1143.8 1160 1200 1240 12777

32.6 65.2 978 130.4 163 0 168 5 1956 228 2 260.8 293 4 326 0 358.6 391 2 423.8 456.4 470.6 489.0 521.6 554.2 566.3 586.8 6194 652.0 6846 717.2 749.8 782.2 7824 815.0 847.6 880.2 912.8 932.2 945.4 978.0 1010.6 1041.3

0 0 0 0 0 0 5 19 29 36 46 53 58 62 65 63 53 46 39 37 36 34 32 31 29 28 27 27 26 25 24 23 22 22 21 21 20

MOMENT (kg-mm)

0 0 0 0 0 0 978 4336 7563 10562 14996 19006 22690 26276 29666 29647 25917 23994 21614 20955 21125 21060 20864 21223 20799 20994 21120 21125 21190 21190 21125 20994 20508 20799 20538 21223 20827

LITERS

40 80 120 160 200 206 7 240 280 320 360 400 440 480 520 560 577.4 600 640 680 694.9 720 760 800 840 880 920 959.8 960 1000 1040 1080 1120 1143.8 1160 1200 1240 1277.7

(kg)

312 62.3 93 5 1246 1558 161.0 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 4498 467 4 498.6 529.7 541.3 560.9 592 0 623.2 6544 685.5 716.7 747.7 747.8 779.0 810.2 841.3 872.5 891.0 903.6 9348 966.0 995 3

(mm)

0 0 0 0 0 0 5 19 29 36 46 53 58 62 65 63 53 46 39 37 36 34 32 31 29 28 27 27 26 25 24 23 22 22 21 21 20

(kg-mm)

0 0 0 0 0 0 935 4144 7229 10096 14334 18166 21687 25115 28356 28337 24772 22934 20659 20029 20192 20129 19942 20285 19880 20067 20187 20192 20254 20254 20192 20067 19602 19880 19631 20285 19907

Most critical fuel amount for right side most lateral C.G. condition. Weights given are nominal weights at 15 C.

412-FMS-63-5-12-2

Rev.1

27


Table 5-13. USABLE FUEL LOADINGTABLE WITHSELF-SEALINGTANKS AND 32.6 GAL AUX FUEL (English)

Jet A, A-1,JP-5, JP-8 (6.8 Lbs/Gal)

U.S. WEIGHTLONGCG MOMENT GALLON (lbs.) (in.) (in-lb)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 168.9 170 180 190 199.9 200 210 220 230 240 250 260 269.9 270 280 290 300 310 318.5 320 330 340 350 353.9

68 136 204 272 340 371.3 408 476 544 612 680 748 816 884 952 1020 1088 1148.5 1156 1224 1292 1359.3 1360 1428 1496 1564 1632 1700 1768 1835.3 1836 1904 1972 2040 2108 2165.8 2176 2244 2312 2380 2406.5

139.5 139.7 139.8 139.9 139.9 139.9 143.1 147.0 149.8 151.8 153.3 154.5 155.6 156.4 157.2 157.9 158.5 159.2 158.7 155.1 151.9 149.0 149.0 150.2 151.4 152.5 153.4 154.3 155.1 155.9 155.9 154.4 153.0 151.7 150.5 149.5 149.6 150.4 151.2 151.8 152.1

9486 18999 28519 38053 47566 51942

58385 69972

81491 92902 104244 115566 126970 138258 149654 161058

172448 182844

183457 189842 196255 202539 202640

214486 226494 238510 250349 262310 274217 286126 286232

293978 301716

309468 317254 323787 325530 337498 349574 361284 366032


U.S. WEIGHT LONGCG MOMENT GALLON (lbs.) (in.) (in-lb)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 168.9 170 180 190 199.9 200 210 220 230 240 250 260 269.9 270 280 290 300 310 318.5 320 330 340 350 353.9

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 1040 1097.9 1105 1170 1235 1299.4 1300 1365 1430 1495 1560 1625 1690 1754.4 1755 1820 1885 1950 2015 2070.3 2080 2145 2210 2275 2300.4

139.5 139.7 139.8 139.9 139.9 139.9 143.1 147.0 149.8 151.8 153.3 154.5 155.6 156.4 157.2 157.9 158.5 159.2 158.7 155.1 151.9 149.0 149.0 150.2 151.4 152.5 153.4 154.3 155.1 155.9 155.9 154.4 153.0 151.7 150.5 149.5 149.6 150.4 151.2 151.8 152.1

9068 18161 27261 36374 45468 49651 55809 66885 77896 88803 99645 110468 121368 132158 143052 153953 164840 174778 175364 181467 187597 193603 193700 205023 216502 227988 239304 250738 262119 273503 273605 281008 288405 295815 303258 309502 311168 322608 334152 345345 349883

Most critical amount for most forward C.G. condition at a weight empty below 6580 pounds has no fuel. * Most critical amount for most forward C.G. condition at a weight empty of 6580 pounds or greater. ** Most critical fuel amount for most aft C.G. condition at a weight empty below 6550 pounds. ***Most critical fuel amount for most aft C.G. condition at a weitght empty of 6550 pounds or geater. Weights given are nominal weights at 15 C. 412-FMS-63-5-13-1

28

Rev. 1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-13M. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 123 LITERS AUX FUEL (Metric) Jet A, A-1, JP-5, JP-8 (.815kg/)

LITERS

40 80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 639.1 640 680 720 756.6 760 800 840 880 920 960 1000 1021.5 1040 1080 1120 1160 1200 1205.5 1240 1280 1320 1339.4

*

WEIGHT LONG CG (kg) (mm)

32.6 65.2 97.8 130.4 1630 168.5 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 520.9 521.6 554.2 586.8 616.6 619.4 652.0 684.6 717.2 749.8 782.4 815.0 832.5 847.6 880.2 912.8 945.4 978.0 982.5 1010.6 1043.2 1075.8 1091.6

3543 3548 3551 3553 3553 3553 3675 3762 3831 3877 3913 3942 3967 3988 4006 4023 4043 4041 3943 3855 3784 3787 3830 3851 3879 3904 3927 3949 3959 3940 3902 3866 3833 3802 3797 3816 3836 3865 3864

Jet B. JP-4 (.779 kg/)

MOMENT (kg-mm)

115502 231330 347288 463311 579139 598540 718830 858488 999125 1137512 1275638 1413601 1551890 1690114 1828338 1967247 2105863 2107786 2185211 2262114 2333324 2345668 2497160 2636395 2782019 2927219 3072485 3218435 3295957 3339544 3434540 3528885 3623718 3718356 3730486 3856450 4001715 4157967 4217985

LITERS

40

-

80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 639.1 640 680 720 756.6 760 800 840 880 920 960 1000 1021.5 1040 1080 1120 1160 1200 1205.5 1240 1280 1320 1339.4

WEIGHT LONG CG (kg) (mm)

31.2 62.3 93.5 124.6 1558 161 0 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 497.9 498.6 529.7 560.9 589.4 592.0 623.2 654.4 685.5 716.7 747.8 779.0 795.7 810.2 841.3 872.5 903.6 934.8 939.1 966.0 997.1 1028.3 1043.4

3543 3548 3551 3553 3553 3553 3675 3762 3831 3877 3913 3942 3967 3988 4006 4023 4043 4041 3943 3855 3784 3787 3830 3851 3879 3904 3927 3949 3959 3940 3902 3866 3833 3802 3797 3816 3836 3865 3864

MOMENT (kg-mm)

110400 221111 331947 442846 553557 572102 687078 820567 954992 1087266 1219291 1351160 1483341 1615459 1747577 1880350 2012844 2014681 2088686 2162192 2230257 2242055 2386856 2519940 2659132 2797919 2936768 3076271 3150368 3192030 3282831 3373008 3463652 3554110 3565704 3686103 3824952 3974302 4031669

Most critical amount for most forward C.G. condition at a weight empty below 2984 kilograms has no fu Most critical amount for most forward C.G. condition at a weight empty of 2984 kilograms or greater. ** Most critical fuel amount for most aft C.G. condition at a weight empty below 2971 kilograms. Most critical fuel amount for most aft C.G. condition at a weight empty of 2971 kilograms or greater. Weights given are nominal weights at 15 C. 412-FMS-63-5-13-2

Rev. 1

29


Table 5-14. USABLE FUEL LOADINGTABLE WITHSELF-SEALINGTANKS AND 32.6 GAL AUX FUEL (English)

Jet A, A-1, JP-5, JP-8 (6.8 Lbs/Gal)


U.S. WEIGHT LAT CG MOMENT GALLON (lbs.) (In.) (in-lb)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 168.9 170 180 190 199.9 200 210 220 230 240 250 260 269.9 270 280 290 300 310 318.5 320 330 340 350 353.9

68 136 204 272 340 371.3 408 476 544 612 680 748 816 884 952 1020 1088 1148.5 1156 1224 1292 1359.3 1360 1428 1496 1564 1632 1700 1768 1835.3 1836 1904 1972 2040 2108 2165.8 2176 2244 2312 2380 2406.5

0 0 0 0 0 0 -0.06 -0.04 -0.05 -0.04 -0.04 -0.03 -0.03 -0.03 -0.03 -0.03 -0.03 -0.02 -0 18 -0.36 -0.48 -0.58 -0.58 -0.55 -0.53 -0.51 -0.48 -0.47 -0.44 -0.43 -0.43 -0.41 -0.40 -0.39 -0.37 -0.36 -0.36 -0.35 -0.34 -0.33 -0.33

0 0 0 0 0 0

-24 -19

-27 -24 -27 -22 -24 -27 -29 -31 -33

-23 -208

-441 -620 -788 -789

-785 -793

-798 -783

-799 -778 -789 -789

-781 -789 -796

-780 -780 -783 -785 -786

-785 -794

U.S. WEIGHT LAT CG GALLON (lbs.) (in.)

10 20 30 40 50 54.6 60 70 80 90 100 110 120 130 140 150 160 168.9 170 180 190 199.9 200 210 220 230 240 250 260 269.9 270 280 290 300 310 318.5 320 330 340 350 353.9


30

Rev. 1

65 130 195 260 325 354.9 390 455 520 585 650 715 780 845 910 975 1040 1097.9 1105 1170 1235 1299.4 1300 1365 1430 1495 1560 1625 1690 1754.4 1755 1820 1885 1950 2015 2070.3 2080 2145 2210 2275 2300.4

0 0 0 0 0 0 -0.06 -0.04 -0.05 -0.04 -0.04 -0.03 -0.03 -0.03 -0.03 -0.03 -0.03 -0.02 -0.18 -0.36 -0.48 -0.58 -0.58 -0.55 -0.53 -0.51 -0.48 -0.47 -0.44 -0.43 -0.43 -0.41 -0.40 -0.39 -0.37 -0.36 -0.36 -0.35 -0.34 -0.33 -0.33

MOMENT (in-lb)

0 0 0 0 0

0

-23

-18 -26 -23 -26 -21 -23 -25 -27 -29 -31

-22 -199

-421 -593

-754

-754

-751 -758

-762 -749 -764

-744 -754 -755

-746 -754 -761 -746

-745 -749 -751 -751 -751

-759

412-FMS-63-5-14-1

BHT-412-FMS-63.2 & 63.3 & 63.4

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O Table 5-14M. USABLE FUEL LOADING TABLE WITH SELF-SEALING TANKS AND 123 LITERS AUX FUEL (Metric)

LITERS

WEIGHT (kg)

40

32.6

80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 639.1 640 680 720 756.6 760 800 840 880 920 960 1000 1021.5 1040 1080 1120 1160 1200 1205.5 1240 1280 1320 1339.4

65.2 97.8 130.4 163.0 168.5 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 520.9 521.6 554.2 586.8 616.6 619.4 652.0 684.6 717.2 749.8 782.4 815.0 832.5 847.6 880.2 912.8 945.4 978.0 982.5 1010.6 1043.2 1075.8 1091.6

LATCG (mm)

MOMENT (kg-mm)

0

0

0 0 0 0 0 -1 -1 -1 -1 -1 -1 -1 -1 -1 -1 -1 -1 -9 -12 -15 -15 -14 -13 -13 -12 -12 -11 -11 -11 -10 -10 -10 -9 -9 -9 -9 -8 -8

0 0 0 0 0 -196 -228 -261 -293 -326 -359 -391 -424 -456 -489 -521 -522 -4988 -7042 -9249 -9291 -9128 -8900 -9324 -8998 -9389 -8965 -9158 -9324 -8802 -9128 -9454 -8802 -8842 -9095 -9389 -8606 -8733

LITERS

WEIGHT (kg)

LAT CG (mm)

40

31.2

0

0

62.3 93.5 124.6 155.8 161.0 187.0 218.1 249.3 280.4 311.6 342.8 373.9 405.1 436.2 467.4 497.9 498.6 529.7 560.9 589.4 592.0 623.2 654.4 685.5 716.7

0 0 0 0 0 -1 -1

0 0 0 0 0 -187 -218

-1 -1 -1 -1 -1 -1 -1 -1 -1 -1 -9 -12 -15 -15 -14 -13 -13 -12 -12 -11 -11 -11 -10 -10 -10 -9 -9 -9 -9 -8 -8

-249 -280 -312 -343 -374 -405 -436 -467 -498 -499 -4767 -6731 -8841 -8881 -8725 -8507 -8912 -8600 -8974 -8569 -8753 -8912 -8413 -8725 -9036 -8413 -8452 -8694 -8974 -8226 -8347

80 120 160 200 206.7 240 280 320 360 400 440 480 520 560 600 639.1 640 680 720 756.6 760 800 840 880 920

960 .

1000 1021.5 1040 1080 1120 1160 1200 1205.5 1240 1280 1320 1339.4

747.8

779.0 795.7 810.2 841.3 872.5 903.6 934.8 939.1 966.0 997.1 1028.3 1043.4

MOMENT (kg-mm)

* Most critical fuel amount for left side most lateral C.G. condition.

Weights given are nominal weights at 15 C.

412-FMS-63-5-14-2

Rev. 1

31/32


MODELS


TEN CELL

SELF SEALING FUEL

412-899-377

S/N 33108 -

33213

AND

S/N 36020 -

36086

S/N 36087 AND SUB CERTIFIED 22 JUNE 1998

This supplement shall be attached to Model 412 or 412EP Flight Manual (BHT-412-FM-2,BHT-412-FM-3and BHT-412-FM-4),when SELF SEALING FUEL CELLS are installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, or other applicable supplements, consult basic Flight Manual.

COPYRIGHT NOTICE COPYRIGHT 1998 BELL HELICOPTER TEXTRON INC ANDBELLHELICOPTER TEXTRON INC. ADIVISION OFTEXTRON CANADA LTD. ALLRIGHTS RESERVED

22 JUNE 1998

REVISION1 - 02 JULY 1998


NOTICEPAGE

PROPRIETARYRIGHTS NOTICE



NP


LOG OF REVISIONS


0 ..............

22 JUN 98

Revision ..........

..............

02 JUL 98

LOG OF PAGES

REVISION

REVISION NO.

PAGE

FLIGHT MANUAL

Title .................................... NP ......................................

1 0

NO.

PAGE

1 0 1 0

A - B .................................. i/ii ....................................... 1 ........................................ 2 - 32 ..................................

NOTE


A



Original...........0 Revision ..........

.............. ..............

22 JUN 98 02 JUL 98

APPROVED:

ROTORCRAFTCERTIFICATIONOFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH,TX 76193-0170

B

Rev.1

DATE: JUL 0 2

1998

BHT-412-FMS-65.2 & 65.3 & 65.4


Ten Cell-Self Sealing Fuel Cells (412-899-377) replaces have self-sealing capability against ballistic damage of self-sealing fuel cells are identical to standard different since self-sealing cells have greater wall thickness of cell, total and usable fuel capacities are All other components system.

including

fuel quantity

standard fuel cells with cells which by small objects. Exterior geometry cells. Interior geometry is slightly thickness. Due to increased wall less than those of basic helicopter.

gauging

system

are identical

to standard

i/ii

T N I * Section 1 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

1-6.

BHT-412-FMS-65.2 & 65.3 & 65.4

WEIGHT AND CENTER OF

1-20.

INSTRUMENT

MARKINGS AND PLACARDS

GRAVITY

Refer to figure 1-1.

Actual weight change shall be determined after installation and ballast readjusted, if necessary, to retain gross weight CG within allowable limits.

BASIC FUEL CAP 2121 LBS

WITH AUX FUEL KIT

412-706-007 3236

LBS

412-706-009 2343 LBS

Location: Instrument panel

Figure 1-1.

Instrument Markings and Placards

Rev.

1

T N I Section 2 I R P N O R I O T F C U T O OD N R E P E E R R C S Section 4 N G O N I R W O E I F V NLY O BHT-412-FMS-65.2 & 65.3 & 65.4

FAA APPROVED


Section 3



2

T N I Section5 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-65.2 & 65.3 & 65.4

5-7. FUEL LOADING

Fuel loading tables lists usable fuel quantities in 10 gallon (40 liter) increments, with weights and moments in both english and metric units for balance computation. Critical fuel loading for computing most forward and aft CGs are denoted.

5-7-A. BASICSYSTEM- SELF SEALINGFUEL CELL.

5-7-C. BASIC SYSTEMWITHONE SEATTYPEAUXILIARYFUELTANK (LEFT OR RIGHT). Refer to BHT-412-FMS-25.2/25.3/25.4.


Total capacity: 326.5 U.S. gallons (1235.9 liters). Usable fuel: 317.3 U.S. gallons (1201.1 liters). Tables 5-1 - Provides longitudinal CG data for approved fuels. * Tables 5-2 - Provides lateral CG data for approved fuels. ·

5-7-B. BASIC SYSTEMWITH BOTH SEATTYPEAUXILIARYTANKS (16.3 GAL /61.7 LITREEACH). Refer to BHT-412-FMS-25.2/25.3/25.4.


Table 5-3 - Provides longitudinal CG data for approved fuels. Table 5-4 - Provides lateral CG data for approved fuels.

* *

Table 5-5 - Provides longitudinal CG data for left side installation. Table 5-6 - Provides lateral CG data for left side installation. Table 5-7 - Provides longitudinal CG data for right side installation. Table 5-8 - Provides lateral CG data for right side installation.

5-7-D. BASICSYSTEMWITH BOTH LONGRANGEAUXILIARYFUEL TANKS(81.7 GAL /309.3 LITRE EACH). Refer to BHT-412-FMS-17.2/17.3/17.4.

Total capacity: 489.9 U.S. gallons (1854.4 liters). Usable fuel: 480.7 U.S. gallons (1819.6 liters). * ·

Table 5-9 - Provides longitundinal CG data for approved fuels. Table 5-10- Provides lateral CG data for approved fuels.


Usable fuel: 399.0 U.S. gallons (1510.4 liters).

5-7-E. BASICSYSTEMWITHONE LONGRANGEAUXILIARYFUEL TANK (LEFT OR RIGHT). Refer to BHT-412-FMS-17.2/17.3/17.4.

Total capacity: 408.2 U.S. gallons (1545.2 liters).

* Table 5-11 - Provides longitudinal CG data for left side installation. * Table 5-12 - Provides lateral CG data for left side installation. * Table 5-13 - Provides longitudinal CG data for right side Installation. * Table 5-14 - Provides lateral CG data for right side installation.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412FMS-65.2 & 65.3 & 65.4

Table 5-1.

Usable fuel loading - longitudinal - basic helicopter (English)

USABLE FUEL LOADING TABLE - LONGITUDINAL CG (ENGLISH) Basic System - 317.3 US Gal Jet B, JP-4 (6.5 lb./Gal) Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) CG Weight Quantity Moment CG Weight Quantity (inches) (USGal) (lb.) (in-b) (lb.) (inches) (USGal) 138.6 65 9424 10 138.6 10 68 139.1 130 18921 20 139.1 20 136 139.4 195 30 28442 139.4 204 30 260 139.6 40 139.6 37962 40 272 139.7 325 50 47483 139.7 340 50 139.7 60 390 57005 408 139.7 60 139.7 407 *62.6 59486 139.7 426 62.6 143.6 455 70 68334 143.6 476 70 147.7 80 520 80322 544 147.7 80 150.8 90 585 92289 612 150.8 90 153.3 650 100 104245 153.3 680 100 715 155.3 110 155.3 116199 110 748 157.0 780 128152 120 157.0 120 816 158.5 130 845 140108 884 158.5 130 159.6 903 *138.9 150706 159.6 944 **138.9 910 159.1 140 159.1 151425 952 140 154.7 975 150 157776 1020 154.7 150 150.9 1040 160 164126 150.9 1088 160 147.5 1105 170481 170 147.5 170 1156 147.4 *170.4 1107 170708 1158 147.4 *170.4 148.9 180 1170 182237 1224 148.9 180 150.3 1235 190 194191 150.3 1292 190 1300 151.6 200 151.6 206144 200 1360 152.7 1365 210 218097 152.7 1428 210 153.8 1430 220 230048 153.8 1496 220 154.7 1495 230 241999 154.7 1564 230 155.3 *237.8 1546 155.3 251064 1617 **237.8 1560 154.9 240 154.9 252760 1632 240 153.3 1625 250 260527 1700 153.3 250 151.8 1690 260 268295 151.8 1768 260 1755 150.4 270 150.4 276062 1836 270 149.1 1820 280 283824 149.1 1904 280 148.3 *286.4 1948 148.3 288808 1948 286.4 1885 148.6 290 148.6 293083 1972 290 1950 149.5 300 149.5 305032 300 2040 150.4 2015 310 316981 150.4 2108 310 151.0 2062 317.3 325703 151.0 2158 **317.3

Moment (In-b) 9009 18086 27187 36288 45388 54490 56862 65319 76778 88217 99646 111073 122498 133926 144057 144744 150815 156885 162960 163177 174197 185624 197050 208475 219899 231322 239988 241609 249033 256458 263883 271302 288808 280153 291575 302996 311334

412-FMS-65-5-1-1

5

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412FMS-65.2& 65.3 & 65.4

Table 5-1. Usablefuel loading- longitudinal- basic helicopter(Metric) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (METRIC) Basic System - 1201.1 litres Jet A, A-1, JP-5, JP-8 (0.815 kgA) Jet B, JP-4 (0.799 kgA) Quantity Weight CG Moment Quantity Weight CG 40 80 120 160 200 *237.0 240 280 320 360 400 440 480 520 525.7 560 600 640 644.9 680 720 760 800 840 880 **900.2 920 960 1000 1040 1080 1084.2 1120 1160 1200 **1201.1

32.6 65.2 97.8 130.4 163.0 193.1 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 428.4 456.4 489.0 521.6 525.6 554.2 586.8 619.4 652.0 684.6 717.2 733.7 749.8 782.4 815.0 847.6 880.2 883.6 912.8 945.4 978.0 978.9

3520 3534 3541 3545 3547 3548 3560 3690 3788 3864 3924 3973 4014 4048 4053 3950 3845 3753 3742 3779 3818 3852 3882 3910 3935 3943 3920 3878 3839 3802 3769 3766 3788 3811 3833 3833

114758 230430 346339 462247 578156 685353 696327 842099 987966 1133564 1279108 1424637 1570171 1715733 1736323 1802684 1879979 1957351 1966775 2094599 2240135 2385658 2531170 2676669 2822156 2892574 2939298 3033864 3128430 3222967 3317451 3327431 3457537 3603004 3748467 3752508

40 80 120 160 200 *237.0 240 280 320 360 400 440 480 520 **525.7 560 600 640 644.9 680 720 760 800 840 880 **900.2 920 960 1000 1040 1080 *1084.2 1120 1160 1200 **1201.1

32.0 63.9 95.9 127.8 159.8 189.3 191.8 223.7 255.7 287.6 319.6 351.6 383.5 415.5 420.0 447.4 479.4 511.4 515.3 543.3 575.3 607.2 639.2 671.2 703.1 719.3 735.1 767.0 799.0 831.0 862.9 866.3 894.9 926.8 958.8 959.7

3520 3534 3541 3545 3547 3548 3560 3690 3788 3864 3924 3973 4014 4048 4053 3950 3845 3753 3742 3779 3818 3852 3882 3910 3935 3943 3920 3878 3839 3802 3769 3766 3788 3811 3833 3833

Moment

112505 225907 339540 453172 566805 671898 682657 825567 968570 1111310 1253997 1396668 1539345 1682050 1702236 1767294 1843072 1918925 1928163 2053478 2196156 2338823 2481478 2624121 2766752 2835788 2881594 2974303 3067013 3159694 3252323 3262107 3389659 3532270 3674877 3678839

NOTE This table is invalid with auxiliary fuel tank(s) installed

412-FMS65-5-1-2


Table 5-2.

Usable fuel loading - lateral - basic helicopter (English)

USABLE FUEL LOADING TABLE - LATERAL CG (ENGLISH) Basic System - 317.3 US Gal Jet A, A-1, JP-5, JP-8 (6.8 b./Gal) Jet B, JP-4 (6.5 lb./Gal) Quantity Weight CG Moment Quantity Weight CG (USGal) (lb ) (Inches) (-b ) (USGal) (lb) (inches) 10 68 0.0 0 10 65 0.0 20 136 0.0 0 20 130 0.0 30 204 0.0 0 30 195 0.0 40 272 0.0 0 40 260 0.0 50 340 0.0 0 50 325 0.0 60 408 0.0 0 60 390 0.0 62.6 426 0.0 0 62.6 407 0.0 70 476 0.0 -13 70 455 0.0 80 544 0.0 -13 80 520 0.0 90 612 0.0 -13 90 585 0.0 100 680 0.0 -13 100 650 0.0 110 748 0.0 -13 110 715 0.0 120 816 0.0 -13 120 780 0.0 130 884 0.0 -13 130 845 0.0 138.9 944 0.0 -13 138.9 903 0.0 140 952 -0.2 -171 140 910 -0.2 150 1020 -0.4 -407 150 975 -0.4 160 1088 -0.5 -540 160 1040 -0.5 170 1156 -0.6 -672 170 1105 -0.6 170.4 1158 -0.6 -677 *170.4 1107 -0.6 180 1224 -0.6 -677 180 1170 -0.6 190 1292 -0.5 -677 190 1235 -0.5 200 1360 -0.5 -677 200 1300 -0.5 210 1428 -0.5 -677 210 1365 -0.5 220 1496 -0.5 -677 220 1430 -0.5 230 1564 -0.4 -677 230 1495 -0.4 237.8 1617 -0.4 -677 237.8 1546 -0.4 240 1632 -0.4 -677 240 1560 -0.4 250 1700 -0.4 -677 250 1625 -0.4 260 1768 -0.4 -677 260 1690 -0.4 270 1836 -0.4 -677 270 1755 -0.4 280 1904 -0.4 -677 280 1820 -0.4 286.4 1948 -0.3 -677 286.4 1862 -0.3 290 1972 -0.3 -677 290 1885 -0.3 300 2040 -0.3 -678 300 1950 -0.3 310 2108 -0.3 -688 310 2015 -0.3 317.3 2158 -0.3 -696 317.3 2062 -0.3

Moment (in-b) 0 0 0 0 0 0 0 -12 -13 -13 -13 -13 -13 -13 -13 -164 -389 -516 -642 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -648 -658 -665


412-FMS-65-5-2-1

7

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412FMS-65.2 & 65.3 & 65.4 Table 5-2.

120 160 200 237.0 240 280 320 360 400 440 480 520 525.7 560 600 640 *644.9 680 720 760 800 840 880 900.2 920 960 1000 1040 1080 1084.2 1120 1160 1200 1201.1

Usable fuel loading - lateral - basic helicopter (Metric)

97.8 130.4 163.0 193.2 195.6 228.2 260.8 293.4 326.0 358.6 391.2 423.8 428.4 456.4 489.0 521.6 525.6 554.2 586.8 619.4 652.0 684.6 717.2 733.7 749.8 782.4 815.0 847.6 880.2 883.6 912.8 945.4 978.0 978.9

0 0 0 0 0 -1 -1 -1 0 0 0 0 0 -10 -12 -15 -15 -14 -13 -13 -12 -11 -11 -11 -10 -10 -10 -9 -9 -9 -9 -8 -8 -8

0 0 0 0 -24 -150 -152 -152 -152 -152 -152 -152 -152 -4388 -5989 -7593 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7796 -7803 -7885 -8014 -8018

120 160 200 237.0 240 280 320 360 400 440 480 520 526 560 600 640 *644.9 680 720 760 800 840 880 900.2 920 960 1000 1040 1080 1084.2 1120 1160 1200 1201.1

95.9 127.8 159.8 189.4 191.8 223.7 255.7 287.6 319.6 351.6 383.5 415.5 420.0 447.4 479.4 511.4 515.3 543.3 575.3 607.2 639.2 671.2 703.1 719.3 735.1 767.0 799.0 831.0 862.9 866.3 894.9 926.8 958.8 959.7


0 0 0 0 0 -1 -1 -1 0 0 0 0 0 -10 -12 -15 -15 -14 -13 -13 -12 -11 -11 -11 -10 -10 -10 -9 -9 -9 -9 -8 -8 -8

0 0 0 0 -24 -147 -149 -149 -149 -149 -149 -149 -149 -4301 -5872 -7444 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7649 -7730 -7857 -7860


Table 5-3.

Usable fuel loading - longitudinal - w/32.6 gal aux fuel In LH and RH position (English) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (ENGLISH) Basic with LH and RH 16.3 US Gal Aux Tank - 349.9 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) ity Weight CG Moment Quantity Weight CG

10 20 30 40 50 60

70 80 90 100 110 120 130 140 150 160 170 171 .5 180 190 200 203. 0 210 220 230 240 250 260 270 .4 280 290 300 310 0

9

68 136 204 272 340 408 426 476 544 612 680 748 816 884 952 1020 1088 1156 1166 1224 1292 1360 1380 1428 1496 1564 1632 1700 1768 1836 1839 1904 1972 2040 2108 2169 2176 2244 2312 2379

138.6 139.1 139.4 139.6 139.7 139.7 139.7 143.2 146.3 148.9 150.7 152.2 153.4 154.4 155.3 156.1 156.7 157.5 157.7 154.7 151.4 148.5 147.7 148.7 149.9 151.0 152.1 153.0 153.9 154.6 154.6 153.2 151.8 150.6 149.4 148.4 148.5 149.3 150.1 150.8

9424 18921 28442 37962 47483 57005 59486 68162 79611 91140 102497 113831 125151 136497 147835 159172 170527 182123 183874 189293 195642 201997 203876 212296 224251 236204 248157 260109 272059 283909 284232 291675 299443 307211 314974 321976 323144 335093 347042 358871

10 20 30 40 50 60 *62.6 70 80 90 100 110 120 130 140 150 160 170 171.5 180.0 190 200 *203.0 210 220 230 240 250 260 270 270.4 280 290 300 310 *319.0 320 330 340 **349.9

65 130 195 260 325 390 407 455 520 585 650 715 780 845 910 975 1040 1105 1115 1170 1235 1300 1319 1365 1430 1495 1560 1625 1690 1755 1758 1820 1885 1950 2015 2074 2080 2145 2210 2274

138.6 139.1 139.4 139.6 139.7 139.7 139.7 143.2 146.3 148.9 150.7 152.2 153.4 154.4 155.3 156.1 156.7 157.5 157.7 154.7 151.4 148.5 147.7 148.7 149.9 151.0 152.1 153.0 153.9 154.6 154.6 153.2 151.8 150.6 149.4 148.4 148.5 149.3 150.1 150.8

Moment

9009 18086 27187 36288 45388 54490 56862 65155 76099 87119 97975 108809 119630 130475 141313 152149 163004 174088 175762 180942 187011 193085 194882 202930 214358 225784 237209 248633 260057 271384 271692 278807 286232 293657 301078 307771 308888 320310 331731 343038

Criticalfuel quantities for mostfoward CGconditon(Includes zero uel) Criticalfuel quanitiesformostat CGcondition

Weightsgivenare nominal weightsat 15"C (59' F)

412-FMS-65-5-3-1

9


Table 5-3. Usable fuel loading- longitudinal- w/123.4 litres aux fuel in LH and RH position(Metric) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (METRIC) Basic with LH and RH 61.7 litre Aux Tank - 1324.5 litres

40 80 120 160 200 *237. 240 280 320 360 400 440 480 520 560 600 640 **649 680 720 760 *768. 800 840 880 920 960 1000

**1023.6

"1324


Table 5-4.

Usable fuel loading - lateral - w/32.6 gal. aux fuel in LH and RH position (English)

USABLE FUEL LOADING TABLE - LATERAL CG (ENGLISH) Basic with LH and RH 16.3 US Gal Aux Tank - 349.9 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Quantity Weight CG Moment Quantity Weight CG (USGal) (lb) (inches) (in-lb) (USGal) (lb) (inches) 10 68 0.0 0 10 65 0.0 20 136 0.0 0 20 130 0.0 30 204 0.0 0 30 195 0.0 40 272 0.0 0 40 260 0.0 50 340 0.0 0 50 325 0.0 60 408 0.0 0 60 390 0.0 62.6 426 0.0 0 62.6 407 0.0 70 476 0.0 -11 70 455 0.0 80 544 0.0 -13 80 520 0.0 90 612 0.0 -13 90 585 0.0 100 680 0.0 -13 100 650 0.0 110 748 0.0 -13 110 715 0.0 120 816 0.0 -13 120 780 0.0 130 884 0.0 -13 130 845 0.0 140 952 0.0 -13 140.0 910 0.0 150 1020 0.0 -13 150 975 0.0 160 1088 0.0 -13 160 1040 0.0 170 1156 0.0 -13 170 1105 0.0 171.5 1166 0.0 -13 171 1115 0.0 180 1224 -0.3 -374 180.0 1170 -0.3 190 1292 -0.4 -504 190 1235 -0.4 200 1360 -0.5 -637 200 1300 -0.5 203.0 1380 -0.5 -677 203.0 1319 -0.5 210 1428 -0.5 -677 210 1365 -0.5 220 1496 -0.5 -677 220 1430 -0.5 230 1564 -0.4 -677 -0.4 1495 230 240 1632 -0.4 -677 240.0 1560 -0.4 250 1700 -0.4 -677 250 1625 -0.4 260 1768 -0.4 -677 260 1690 -0.4 270 1836 -0.4 -677 270 1755 -0.4 270.4 1839 -0.4 -677 270 1758 -0.4 280 1904 -0.4 -677 280 1820 -0.4 290 1972 -0.3 -677 290.0 1885 -0.3 300 2040 -0.3 -677 300 1950 -0.3 310 2108 -0.3 -677 310 2015 -0.3 319.0 2169 -0.3 -677 319 2074 -0.3 320 2176 -0.3 -677 320 2080 -0.3 330 2244 -0.3 -678 -0.3 2145 330 340 2312 -0.3 -685 340 2210 -0.3 349.9 2379 -0.3 -696 350 2274 -0.3

Moment (In-b) 0 0 0 0 0 0 0 -10 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -357 -482 -609 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -648 -655 -665

Criticalfuelquantityfor mostlateralCGcondition

°

Weightsgivenare nominalweightsat 15' C (59 F) Forcalculation purposes,useWeightandMomentvalues.(CGsshownare derivedfromWeight andMomentvaluesandrounded to onedecimalplace)

412-FMS-65-5-4-1

11


Usable fuel loading - lateral - w/123.4 litre aux fuel in LH and RH position (Metric)

USABLE FUEL LOADING TABLE - LATERAL CG (METRIC) Basic with LH and RH 61.7 litre Aux Tank - 1324.5 litres Jet A, A-1, JP-5, JP-8 (0.815 kg/1) Jet B, JP-4 (0.799 kgA) Quantity Weight CG Moment Quantity Weight CG (litres) (kg) (litres) (kg) (mm) (mm-kg) (mm) 40 32.6 0 0 40 32.0 0 80 65.2 0 0 80 63.9 0 120 97.8 0 0 120 95.9 0 160 130.4 0 0 160 127.8 0 200 163.0 0 0 200 159.8 0 237.0 193.1 0 0 237.0 189.3 0 240 195.6 0 -24 240 191.8 0 280 228.2 -1 -147 280 223.7 -1 320 260.8 1 -152 320 255.7 -1 360 293.4 -1 -152 360 287.6 -1 400 326.0 0 -152 400 319.6 0 440 358.6 0 -152 440 351.6 0 480 391.2 0 -152 480 383.5 0 520 423.8 0 -152 520 415.5 0 560 456.4 0 -152 560 447.4 0 600 489.0 0 -152 600 479.4 0 640 521.6 0 -152 640 511.4 0 649.1 529.0 0 -152 649.1 518.6 0 680 554.2 -8 -4256 -8 543.3 680 720 586.8 -10 -5842 720 575.3 -10 760 619.4 -12 -7458 760 607.2 -12 768.3 626.1 -12 -7796 768.3 613.9 -12 800 652.0 -12 -7796 800 639.2 -12 840 684.6 -11 -7796 840 671.2 -11 880 717.2 -11 -7796 880 703.1 -11 920 749.8 -10 -7796 920 735.1 -10 960 782.4 -10 -7796 960 767.0 -10 1000 815.0 -10 -7796 1000 799.0 -10 1023.6 834.3 -9 -7796 1023.6 817.9 -9 1040 847.6 -9 -7796 1040 831.0 -9 1080 880.2 -9 -7796 1080 862.9 -9 1120 912.8 -9 -7796 1120 894.9 -9 1160 945.4 -8 -7796 1160 926.8 -8 1200 978.0 -8 -7796 1200 958.8 -8 1207.6 984.2 -8 -7796 1207.6 964.9 -8 1240 1010.6 -8 -7800 1240 990.8 -8 1280 1043.2 -8 -7875 1280 1022.7 -8 1320 1075.8 -7 -8003 -7 1054.7 1320 1324.5 1079.5 -7 -8018 1324.5 1058.3 -7

Moment (mm-kg) 0 0 0 0 0 0

-24 -144 -149 -149 -149 -149 -149 -149 -149 -149 -149 -149 -4172 -5727 -7312 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7643 -7647 -7720 -7846 -7860

Criticalluel quantityformostlateralCGcondition

°

°

Weightsgivenare nominal weightsat 15 C. (59 F) Forcalculation purposes,useWeightandMomentvalues.(CGsshownarederivedfromWeight andMomentvaluesandroundedtowholenumbers)

412-FMS-65-5-4-2


Table 5-5.

Usable fuel loading - longitudinal - w/16.3 gal aux fuel in LH position (English)

USABLE FUEL LOADING TABLE - LONGITUDINAL CG (ENGLISH) Basic with LH 16.3 US Gal Aux Tank - 333.6 US Gal Jet B, JP-4 (6.5 lb./Gal) Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Quantily Weight CG Moment Quantity Weight CG (inches) (lb) (USGal) (lb.) (Inches) (in-lb) (USGal (lb) 68 138.6 9424 10 65 138.6 10 136 139.1 18921 20 130 139.1 20 30 204 139.4 28442 30 195 139.4 272 139.6 37962 40 260 139.6 40 50 340 139.7 47483 50 325 139.7 408 139.7 57005 60 390 139.7 60 *62.6 426 139.7 59486 *62.6 407 139.7 70 476 143.4 68237 70 455 143.4 544 146.9 79910 80 520 146.9 80 612 149.7 91587 90 585 149.7 90 100 680 151.7 103177 100 650 151.7 110 748 153.4 114749 110 715 153.4 120 816 154.8 126337 120 780 154.8 130 884 156.0 137919 130 845 156.0 140 952 157.0 149506 140 910 157.0 1020 158.0 161115 150 975 158.0 150 **155.2 1055 158.6 167290 ** 155.2 1009 158.6 160 1040 156.6 156.6 170361 160 1088 152.9 170 1105 152.9 176708 170 1156 180 1224 149.6 183061 180 1170 149.6 * 186.7 1269 147.6 187292 186.7 1213 147.6 190 1292 148.1 191289 190 1235 148.1 1360 149.4 203244 200 1300 149.4 200 210 1428 150.7 215198 210 1365 150.7 220 1496 151.8 227151 220 1430 151.8 1564 152.9 239103 230 1495 152.9 230 240 1632 153.8 251054 240 1560 153.8 154.7 250 1625 154.7 263004 250 1700 ** 254. 1 1728 154.9 267648 **254.1 1652 154.9 260 1690 154.0 154.0 272217 1768 260 270 1836 152.5 279985 270 1755 152.5 151.1 280 1820 280 1904 151.1 287753 1972 149.9 295518 290 1885 149.9 290 300 2040 148.7 303279 300 1950 148.7 2059 148.4 305392 '302.7 1968 148.4 * 302.7 310 2108 149.0 314088 310 2015 149.0 320 2176 149.8 326037 320 2080 149.8 2244 150.6 337985 330 2145 150.6 330 "333. 6 2268 150.9 342287 "333.6 2168 150.9

Moment (in-lb) 9009 18086 27187 36288 45388 54490 56862 65227 76385 87547 98625 109686 120763 131834 142910 154007 159909 162845 168912 174985 179029 182850 194277 205704 217129 228554 239978 251401 255840 260208 267633 275058 282481 289899 291919 300232 311653 323074 327186

Criticalluel quantitiesformostforwardCGcondition(includes zeroluel) ** Criticalluel quantities lor mostaft CGcondiion

Weightsgivenare nominal weightsa 15' C (59' F) purposes,useWeightandMomentvalues.(CGsshownarederivedfromWeight Forcalculation andMoment valuesandroundedtoonedecimalplace).

412-FMS-65-5-5-1


Usable fuel loading - longitudinal - w/61.7 litre aux fuel in LH position (Metric)

USABLE FUEL LOADING TABLE - LONGITUDINAL CG (METRIC) Basic with LH 61.7 litre Aux Tank - 1262.8 litres Jet B, JP-4 (0.799 kg/1) Jet A, A-1, JP-5, JP-8 (0.815 kg/) CG Weight Quantity Quantity CG Moment Weight (mm) (kg) (litres) (litres) (mm) (mm-kg) (kg) 40 32.0 3520 40 32.6 3520 114758 63.9 3534 3534 230430 80 65.2 80 120 95.9 3541 3541 346339 120 97.8 160 127.8 3545 130.4 3545 462247 160 3547 578156 200 159.8 163.0 3547 200 193.1 3548 685353 *237.0 189.3 3548 237.0 191.8 3560 3560 696321 240 195.6 240 280 223.7 3679 3679 839499 228.2 280 320 255.7 3766 260.8 3766 982117 320 3829 360 287.6 293.4 3829 1123351 360 3878 1264312 400 319.6 326.0 3878 400 351.6 3919 3919 1405307 440 358.6 440 391.2 3953 1546342 480 383.5 3953 480 520 415.5 3982 520 423.8 3982 1687364 456.4 4007 1828624 560 447.4 4007 560 469.3 4026 4026 1927390 "587.4 478.7 587.4 600 479.4 3991 489.0 3991 1951824 600 3890 640 511.4 521.6 3890 2029103 640 3801 2106436 680 543.3 680 554.2 3801 *706.6 564.6 3747 575.9 3747 2157842 706.6 3761 720 575.3 586.8 3761 2206699 720 3798 2352242 760 607.2 619.4 3798 760 652.0 3831 2497772 800 639.2 3831 800 684.6 3861 2643290 840 671.2 3861 840 3888 2788798 880 703.1 717.2 3888 880 735.1 3913 3913 2934290 920 920 749.8 960 767.0 3935 3935 3079073 782.4 960 768.6 3933 3933 3083642 "961.9 9 784.0 ** 961.9 799.0 3894 3894 3173624 1000 815.0 1000 847.6 3856 3268190 1040 831.0 3856 1040 880.2 3820 3362756 1080 862.9 3820 1080 912.8 3788 3457256 1120 894.9 3788 1120 1145.9 915.6 3767 933.9 3767 3518498 9 1145.9 945.4 3776 3569678 1160 926.8 3776 1160 3799 1200 958.8 978.0 3799 3715147 1200 3820 3860613 1240 990.8 1240 1010.6 3820 1009.0 3832 3832 3943575 1262.8 1029.2 1262.8

Moment (mm-kg) 112505 225907 339540 453172 566805 671898 682651 823018 962837 1101298 1239491 1377718 1515985 1654238 1792725 1889552 1913506 1989268 2065083 2115479 2163377 2306063 2448736 2591397 2734048 2876684 3018625 3023104 3111319 3204029 3296739 3389383 3449423 3499599 3642212 3784822 3866155

zerotuel) CriticalfuelquantitestormostforwardCG condition(includes Critial fuelquantitiesformostaft CG condition

weightsat 15' C.(59' F) Weightsgivenare nominal purposes, use WeightandMomentvalues.(CGs shownare derivedfrom Weight Forcalculation andMomentvaluesandrounded to wholenumbers)

412-FMS-65-5-5-2


Table 5-6.

Usable fuel - lateral - w/16.3 gal aux fuel in LH position (English)

USABLE FUEL LOADING TABLE - LATERAL CG (ENGLISH) Basic with LH 16.3 US Gal Aux Tank - 333.6 US Gal Jet A, A-1, JP-5 JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Quant ity Weight CG Moment Quantity Weight CG l) (lb ) (inches) (in-lb ) (US Gal) (lb) (inches) (USGa 10 68 0.0 0 10 65 0.0 20 136 0.0 0 20 130 0.0 204 0.0 0 30 195 0.0 30 272 0.0 0 40 260 0.0 40 50 340 0.0 0 50 325 0.0 60 408 0.0 0 60 390 0.0 62.6 426 0.0 0 62.6 407 0.0 70 476 -0.3 -127 70 455 -0.3 80 544 -0.8 -427 80 520 -0.8 90 612 -1.1 -691 90 585 -1.1 680 -1.5 -1019 100 650 -1.5 100 110 748 -1.8 -1345 110 715 -1.8 120 816 -2.0 -1670 120 780 -2.0 130 884 -2.3 -1995 130 845 -2.3 140 952 -2.4 -2316 140 910 -2.4 150 1020 -2.6 -2618 150 975 -2.6 155.2 1055 -2.5 -2618 155.2 1009 -2.5 157.8 1073 -2.7 -2906 157.8 1025 -2.7 160 1088 -2.7 -2932 160 1040 -2.7 170 1156 -2.6 -3059 170 1105 -2.6 180 1224 -2.6 -3194 180 1170 -2.6 186.7 1269 -2.6 -3281 186.7 1213 -2.6 1292 -2.5 -3281 190 1235 -2.5 190 200 1360 -2.4 -3281 200 1300 -2.4 210 1428 -2.3 -3281 210 1365 -2.3 1496 -2.2 -3281 220 1430 -2.2 220 230 1564 -2.1 -3281 230 1495 -2.1 240 1632 -2.0 -3281 240 1560 -2.0 250 1700 -1.9 -3281 250 1625 -1.9 254. 1728 -1.9 -3281 254.1 1652 -1.9 260 1768 -1.9 -3281 260 1690 -1.9 270 1836 -1.8 -3281 270 1755 -1.8 280 1904 -1.7 -3281 280 1820 -1.7 290 1972 -1.7 -3281 290 1885 -1.7 300 2040 -1.6 -3281 300 1950 -1.6 302.7 2059 -1.6 -3281 302.7 1968 -1.6 310 2108 -1.6 -3281 310 2015 -1.6 320 2176 -1.5 -3286 320 2080 -1.5 330 2244 -1.5 -3297 330 2145 -1.5 333.6 2268 -1.5 -3301 333.6 2168 -1.5

Moment (in-lb) 0 0 0 0 0 0 0 -122 -409 -660 -974 -1286 -1596 -1907 -2213 -2502 -2502 -2778 -2803 -2924 -3053 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3137 -3141 -3151 -3155

Critical fuelquantity lor most lateral CG condition

Weights given are nominal weights at 15 C (59° F)

For calculation purposes, use Weight and Moment values, (CGs shown are derived Irom Weight and Moment values and rounded to one decimal place)

412-FMS-65-5-6-1

15

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412FMS-65.2

& 65.3 & 65.4

Table 5-6. Usablefuel - lateral - w/61.7 litre aux fuel in LH position (Metric) USABLE FUEL LOADING TABLE - LATERAL CG (METRIC) Basic with LH 61.7 litre Aux Tank - 1262.8 litres Jet B, JP-4 (0.799 kg/l) Jet A, A-1, JP-5, JP-8 (0.815 kgA) CG Weight Quantity Moment CG Weight Quantity (mm) (kg) (lires) (mm-kg) (kg) (mm) (litres) 0 40 32.0 0 0 32.6 40 0 63.9 80 0 0 80 65.2 0 0 120 95.9 97.8 0 120 127.8 0 0 160 130.4 0 160 159.8 0 0 0 200 163.0 200 0 189.3 237.0 0 0 193.1 237.0 0 191.8 240 -33 0 195.6 240 -13 223.7 280 -2927 -13 228.2 280 -23 320 255.7 -5895 -23 260 8 320 -34 287.6 360 -9903 -34 293.4 360 -43 400 319.6 -13871 -43 400 326.0 -50 351.6 440 -17833 -50 358.6 440 -56 383.5 480 -56 -21788 391.2 480 -61 520 415.5 -25718 -61 423.8 520 -65 560 447.4 -65 -29468 456.4 560 -63 469.3 587.4 -63 -30162 478.7 4 587.4 -69 477.2 597.2 -69 -33480 486.7 597.2 -69 479.4 600 -33580 -69 489.0 600 -67 511.4 640 -35106 -67 521.6 640 -66 543.3 680 -36744 -66 554.2 680 -66 706.6 564.6 -37806 -66 5 575.9 706.6 -64 575.3 720 -37806 -64 586.8 720 -61 607.2 760 -37806 -61 619.4 760 -58 639.2 800 -58 -37806 652.0 800 840 671.2 -55 -55 -37806 684.6 840 -53 880 703.1 -53 -37806 717.2 880 -50 920 735.1 749.8 -50 -37806 920 -48 960 767.0 -37806 -48 782.4 960 -48 768.6 961.9 -37806 784.0 -48 961.9 9 -46 799.0 -37806 1000 -46 815.0 100( 831.0 -45 1040 -37806 -45 847.6 1040 -43 862.9 1080 -37806 -43 880.2 1080 -41 894.9 1120 -37806 -41 912.8 1120 -40 915.6 1145.9 -40 -37806 933.9 1145.9 -40 926.8 -37806 1160 -40 945.4 1160 -39 1200 958.8 -37828 -39 978.0 1200 -38 1240 990.8 1010.6 -38 -37953 1240 -37 -38028 1262.8 1009.0 1262.8 1029.2 -37

Moment (mm-kg) 0 0 0 0 0 0 -32 -2869 -5779 -9708 -13599 -17483 -21361 -25213 -28890 -29570 -32823 -32921 -34417 -36023 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37063 -37085 -37208 -37281

Criticalfuelquantityor mostlateralCG condition

weightsat 15-C.(59- F) Weightsgivenare nominal use WeightandMomentvalues.(CGs shownare derivedfrom Weight Forcalculation purposes, to wholenumbers) andMomentvaluesandrounded

412-FMS-65-5-6-2


Table 5-7.

Usable fuel - longitudinal - w/16.3 gal aux fuel in RH position (English)

USABLE FUEL LOADING TABLE - LONGITUDINAL CG (ENGLISH) Basic with RH 16.3 US Gal Aux Tank - 333.6 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Quantity Weight CG Moment Quantity Weight CG (USGal) (lb.) (Inches) (in-b) (USGal) (lb) (Inches) 10 68 138.6 9424 10 65 138.6 20 136 139.1 18921 20 130 139.1 30 204 139.4 28442 30 195 139.4 40 272 139.6 37962 40 260 139.6 50 340 139.7 47483 50 325 139.7 60 408 139.7 57005 60 390 139.7 *62.6 426 139.7 59486 *62.6 407 139.7 70 476 143.4 68237 70 455 143.4 80 544 146.9 79910 80 520 146.9 90 612 149.7 91587 90 585 149.7 100 680 151.7 103177 100 650 151.7 110 748 153.4 114749 110 715 153.4 120 816 154.8 126337 120 780 154.8 130 884 156.0 137919 130 845 156.0 140 952 157.0 149506 140 910 157.0 150 1020 158.0 161115 150 975 158.0 155.2 1055 158.6 167290 155.2 1009 158.6 160 1088 156.6 170361 160 1040 156.6 170 1156 152.9 176708 170 1105 152.9 180 1224 149.6 183061 180 1170 149.6 *186.7 1269 147.6 187292 186.7 1213 147.6 190 1292 148.1 191289 190 1235 148.1 200 1360 149.4 203244 200 1300 149.4 210 1428 150.7 215198 210 1365 150.7 220 1496 151.8 227151 220 1430 151.8 230 1564 152.9 239103 230 1495 152.9 240 1632 153.8 251054 240 1560 153.8 250 1700 154.7 263004 250 1625 154.7 254.1 1728 154.9 267648 254.1 1652 154.9 260 1768 154.0 272217 260 1690 154.0 270 1836 152.5 279985 270 1755 152.5 280 1904 151.1 287753 280 1820 151.1 290 1972 149.9 295518 290 1885 149.9 300 2040 148.7 303279 300 1950 148.7 *302.7 2059 148.4 305392 *302.7 1968 148.4 310 2108 149.0 314088 310 2015 149.0 320 2176 149.8 326037 320 2080 149.8 330 2244 150.6 337985 330 2145 150.6 **333.6 2268 150.9 342287 **333.6 2168 150.9

Moment (in-b) 9009 18086 27187 36288 45388 54490 56862 65227 76385 87547 98625 109686 120763 131834 142910 154007 159909 162845 168912 174985 179029 182850 194277 205704 217129 228554 239978 251401 255840 260208 267633 275058 282481 289899 291919 300232 311653 323074 327186

Criticalfuel quantitieslor most forward CGcondition (includeszerotuel) ** Criticaluelquantities for mostaft CGcondition

°

Weightsgivenarenominalweightsat 15 C (59 F) Forcalculation purposes. useWeightandMomentvalues.(CGsshownarederivedfromWeight andMomentvaluesandroundedtoonedecimalplace).

412-FMS-65-5-7-1

17


Table 5-7. Usable fuel - longitudinal- w/61.7 litre aux fuel in RH position (Metric) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (METRIC) Basic with RH 61.7 litre Aux Tank - 1262.8 litres Jet A, A-1, JP-5, JP-8 (0.815 kgA) Jet B, JP-4 (0.799 kgA) Quantity Weight CG Moment Quantity Weight CG (litres) (kg) (mm) (mm-kg) (litres) (kg) (mm) 40 32.6 3520 114758 40 32.0 3520 80 65.2 3534 230430 80 63.9 3534 120 97.8 3541 346339 120 95.9 3541 160 130.4 3545 462247 160 127.8 3545 200 163.0 3547 578156 200 159.8 3547 * 237.0 193.1 3548 685353 *237.0 189.3 3548 240 195.6 3560 696321 240 191.8 3560 280 228.2 3679 839499 280 223.7 3679 320 260.8 3766 982117 320 255.7 3766 360 293.4 3829 1123351 360 287.6 3829 400 326.0 3878 1264312 400 319.6 3878 440 358.6 3919 1405307 440 351.6 3919 480 391.2 3953 1546342 480 383.5 3953 520 423.8 3982 1687364 520 415.5 3982 560 456.4 4007 1828624 560 447.4 4007 **587.4 478.7 4026 1927390 "587.4 469.3 4026 600 489.0 3991 1951824 600 479.4 3991 640 521.6 3890 2029103 640 511.4 3890 680 554.2 3801 2106436 680 543.3 3801 * 706.6 575.9 3747 2157842 '706.6 564.6 3747 720 586.8 3761 2206699 720 575.3 3761 760 619.4 3798 2352242 760 607.2 3798 800 652.0 3831 2497772 800 639.2 3831 840 684.6 3861 2643290 840 671.2 3861 880 717.2 3888 2788798 880 703.1 3888 920 749.8 3913 2934290 920 735.1 3913 960 782.4 3935 3079073 960 767.0 3935 "961.9 784.0 3933 3083642 **961.9 768.6 3933 1000 815.0 3894 3173624 1000 799.0 3894 1040 847.6 3856 3268190 1040 831.0 3856 1080 880.2 3820 3362756 1080 862.9 3820 1120 912.8 3788 3457256 1120 894.9 3788 1145.9 933.9 3767 3518498 '1145.9 915.6 3767 1160 945.4 3776 3569678 1160 926.8 3776 1200 978.0 3799 3715147 1200 958.8 3799 1240 1010.6 3820 3860613 1240 990.8 3820 ** 1262.8 1029.2 3832 3943575 ** 1262.8 1009.0 3832

Moment (mm-kg) 112505 225907 339540 453172 566805 671898 682651 823018 962837 1101298 1239491 1377718 1515985 1654238 1792725 1889552 1913506 1989268 2065083 2115479 2163377 2306063 2448736 2591397 2734048 2876684 3018625 3023104 3111319 3204029 3296739 3389383 3449423 3499599 3642212 3784822 3866155

CriticalfuelquantitiesformostforwardCGcondition(includes zerofuel) ** Criticaluelquantitestormostaft CGcondition

°

Weightsgivenarenominalweightsat 15 C (59 F) Forcalculation purposes, use WeightandMomentvalues.(CGsshownare derivedIromWeight andMomentvaluesad rounded to wholenumbers)

412-FMS-65-5-7-2


Table 5-8.

Usable fuel - lateral - w/16.3 gal aux fuel in RH position (English)

USABLE FUEL LOADING TABLE - LATERAL CG (ENGLISH) Basic with RH 16.3 US Gal Aux Tank - 333.6 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Quantity Weight Weight CG Moment Quantity CG (USGal) (lb ) (Inches) (in-lb) (USGal) (lb) (inches) 10 68 0.0 0 10 65 0.0 20 136 0.0 0 130 0.0 20 30 204 0.0 0 30 195 0.0 40 272 0.0 0 40 260 0.0 50 340 0.0 0 50 325 0.0 60 408 0.0 0 60 390 0.0 62.6 426 0.0 0 407 62.6 0.0 70 476 0.2 104 70 455 0.2 80 544 0.7 401 80 520 0.7 90 612 1.1 664 585 90 1.1 100 680 1.5 993 100 650 1.5 110 748 1.8 1319 110 715 1.8 120 816 2.0 1644 120 780 2.0 130 884 2.2 1969 130 845 2.2 140 952 2.4 2289 910 140 2.4 150.0 1020 2.5 2592 150.0 975 2.5 155.2 1055 2.5 2592 155.2 1009 2.5 160 1088 2.1 2277 160 1040 2.1 170 1156 1.9 2151 170 1105 1.9 180 1224 1.6 2016 1170 180 1.6 186.7 1269 1.5 1928 186.7 1213 1.5 190 1292 1.5 1928 190 1235 1.5 200 1360 1.4 1928 200 1300 1.4 210 1428 1.4 1928 210 1365 1.4 220 1496 1.3 1928 220 1430 1.3 230 1564 1.2 1928 230 1495 1.2 240 1632 1.2 1928 240 1560 1.2 250 1700 1.1 1928 250 1625 1.1 254.1 1728 1.1 1928 1652 254.1 1.1 260 1768 1.1 1928 260 1690 1.1 270 1836 1.1 1928 270 1755 1.1 280 1904 1.0 1928 280 1820 1.0 290 1972 1.0 1928 290 1885 1.0 300 2040 0.9 1928 300 1950 0.9 302.7 2059 0.9 1928 302.7 1968 0.9 310 2108 0.9 1928 310 2015 0.9 320 2176 0.9 1923 320 2080 0.9 330 2244 0.9 1913 330 2145 0.9 333.6 2268 0.8 1909 333.6 2168 0.8

Moment (in-lb) 0 0 0 0 0 0 0 99 383 635 949 1260 1571 1882 2188 2477 2477 2177 2056 1927 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1843 1838 1828 1825

CriticalfuelquantityformostlateralCGcondition

°

°

Weightsgivenarenominalweightsat 15 C (59 F) Forcalculation purposes, use WeightandMomentvalues,(CGsshownare dervedfromWeight andMomentvaluesandrounded to onedecimalplace)

412-FMS-65-5-8-1

19


Table 5-8. Usable fuel - lateral - w/61.7 litre aux fuel in RH position(Metric) USABLE FUEL LOADING TABLE - LATERAL CG (METRIC) Basic with RH 61.7 litre Aux Tank - 1262.8 litres Jet A, A-1, JP-5, JP-8 (0.815 kg/l) Jet B, JP-4 (0.799 kgA) Quantity Weight CG Moment Quantity Weight CG (litres) (kg) (mm) (mm-kg) (litres) (kg) (mm) 40 32.6 0 0 40 32.0 0 80 65.2 0 0 80 63.9 0

Moment (mm-kg) 0 0

120

97.8

0

0

120

95.9

0

0

160 200 237.0 240 280 320

130.4 163.0 193.1 195.6 228.2 260.8

0 0 0 0 12 21

0 0 0 -16 2631 5591

160 200 237.0 240 280 320

127.8 159.8 189.3 191.8 223.7 255.7

0 0 0 0 12 21

0 0 0 -16 2579 5481

360

293.4

33

9599

360

287.6

33

9410

400 440 480 520 560 567.8 587.4 600 640 680 706.6 720 760 800 840 880 920 960 961.9 1000 1040 1080 1120 1145.9 1160

326.0 358.6 391.2 423.8 456.4 462.8 478.7 489.0 521.6 554.2 575.9 586.8 619.4 652.0 684.6 717.2 749.8 782.4 784.0 815.0 847.6 880.2 912.8 933.9 945.4

42 49 55 60 64 65 62 54 48 42 39 38 36 34 32 31 30 28 28 27 26 25 24 24 23

13567 17529 21484 25414 29164 29864 29858 26439 24914 23276 22214 22214 22214 22214 22214 22214 22214 22214 22214 22214 22214 22214 22214 22214 22214

400 440 480 520 560 * 567.8 587.4 600 640 680 706.6 720 760 800 840 880 920 960 961.9 1000 1040 1080 1120 1145.9 1160

319.6 351.6 383.5 415.5 447.4 453.7 469.3 479.4 511.4 543.3 564.6 575.3 607.2 639.2 671.2 703.1 735.1 767.0 768.6 799.0 831.0 862.9 894.9 915.6 926.8

42 49 55 60 64 65 62 54 48 42 39 38 36 34 32 31 30 28 28 27 26 25 24 24 23

13301 17185 21063 24915 28592 29278 29272 25920 24425 22819 21778 21778 21778 21778 21778 21778 21778 21778 21778 21778 21778 21778 21778 21778 21778

1200

978.0

23

22192

1200

958.8

23

21756

1240 1262.8

1010.6 1029.2

22 21

22066 21992

1240 1262.8

990.8 1009.0

22 21

21633 21560

Criticalfuelquantity for mostlateralCGcondition

Weightsgivenarenominalweightsat 15 C (59 F) Forcalculation purposes. useWeightandMomentvalues.(CGsshownare derivedfromWeight andMomentvaluesandroundedto wholenumbers)

412-FMS-65-5-8-2


Table 5-9.

Usable fuel - longitudinal - w/81.7 gal aux fuel in LH and RH position (English) USABLE FUEL LOADING TABLE - LONGITUDINALCG (ENGLISH)

21


Table 5-9. Usablefuel - longitudinal- w/309.3 litre aux fuel in LH and RH position (Metric)

22


Table 5-10.

Usable fuel - lateral - w/163.4 gal aux fuel in LH and RH position (English) USABLE FUEL LOADINGTABLE - LATERAL CG (ENGLISH) Basic with LH and RH 81.7 US Gal Aux Tank - 480 7 US Gal Jet B, JP-4 (6.5 lb./Gal) A-1. JP-5, JP-8 (6 8 lb /Gal) Jet A, CG Weight Quantity Moment CG Weight Quantity (US Gal)

lb )

(inches)

10 20 30 40 50 60 62.6 70 80 90 100 110 120 130 140 150 160 170 180 190 200 207.9 210 220 230 *239.4 240 250 260 270 280 290 300 310 320 330 340 350 360 370 371.3 380 390 400 410 420.0 430 440 450 460 470 480 480.7

68 136 204 272 340 408 426 476 544 612 680 748 816 884 952 1020 1088 1156 1224 1292 1360 1414 1428 1496 1564 1628 1632 1700 1768 1836 1904 1972 2040 2108 2176 2244 2312 2380 2448 2516 2525 2584 2652 2720 2788 2856 2924 2992 3060 3128 3196 3264 3269

0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 00 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 -0.2 -0.3 -0.4 -0.4 -0.4 -0.4 -0.4 -0.4 -0.4 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2

Moment

(In-lb)

(US Gal)

(lb.)

(inches)

(in-lb)

0 0 0 0 0 0 0 -12 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -257 -419 -554 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -677 -680 -685 -690 -696 -696

10 20 30 40 50 60 62.6 70 80 90 100 110 120 130 140 150 160 170 180 190 200 207.9 210 220 230 *239.4 240 250 260 270 280 290 300 310 320 330 340 350 360 370 371.3 380 390 400 410 420.0 430 440 450 460 470 480 480.7

65 130 195 260 325 390 407 455 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1351 1365 1430 1495 1556 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2405 2414 2470 2535 2600 2665 2730 2795 2860 2925 2990 3055 3120 3125

00 0.0 0.0 0.0 00 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 00 00 0.0 0.0 0.0 -0.2 -0.3 -0.4 -0.4 -0.4 -0.4 -0.4 -0.4 -0.4 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0.3 -0 3 -0.3 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2 -0.2

0 0 0 0 0 0 0 -11 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -13 -246 -401 -529 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -647 -648 -650 -655 -660 -665 -665

Criticalluel quantityformostlateralCG condition °

Weightsgen are nominalweightsat 15 C (59 F) Forcalculationpuposes.use Weight andMomentvalues.(CGsshownare derivedIromWeight

place) valuesandrounded toonedecimal andMoment

412-FMS-65-5-10-1

23


Table 5-10. Usablefuel - lateral - w/618.5 litre aux fuel in LH and RH position(Metric) USABLE FUEL LOADINGTABLE - LATERALCG (METRIC)

0 0 0 0 -1 -1 -1 0


Table 5-11. Usable fuel - longitudinal- w/81.7 gal aux fuel In LH position(English) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (ENGLISH) Basic with LH 81.7 US Gal Aux Tank - 399.0 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Quantity Weight CG Moment Quantity Weight CG (USGal) (lb) (inches) (in-lb) (USGal) (lb) (inches) 10 68 138.6 9424 10 65 138.6 20 136 139.1 18921 20 130 139.1 30 204 139.4 28442 30 195 139.4 40 272 139.6 37962 40 260 139.6 50 340 139.7 47483 50 325 139.7 60 408 139.7 57005 60 390 139.7 62.6 426 139.7 59486 * 62.6 407 139.7 70 476 143.4 68272 70 455 143.4 80 544 146.8 79846 80 520 146.8 90 612 149.0 91159 90 585 149.0 100 680 150.7 102488 100 650 150.7 110 748 152.2 113821 110 715 152.2 120 816 153.4 125153 120 780 153.4 130 884 154.4 136486 130 845 154.4 140 952 155.3 147819 140 910 155.3 150 1020 156.0 159155 150 975 156.0 160 1088 156.7 170503 160 1040 156.7 170 1156 157.3 181857 170 1105 157.3 173.4 1179 157.5 185696 **173.4 1127 157.5 180 1224 155.1 189899 180 1170 155.1 190 1292 151.9 196246 190 1235 151.9 200 1360 149.0 202600 200 1300 149.0 204.9 1393 147.7 205698 204.9 1332 147.7 210 1428 148.1 211519 210 1365 148.1 220 1496 149.0 222872 220 1430 149.0 230 1564 149.8 234226 230 1495 149.8 240 1632 150.5 245579 240 1560 150.5 250 1700 151.1 256932 250 1625 151.1 260 1768 151.7 268285 260 1690 151.7 270 1836 152.3 279637 270 1755 152.3 280 1904 152.8 290988 280 1820 152.8 290 1972 153.3 302340 290 1885 153.3 300 2040 153.8 313691 300 1950 153.8 **304.6 2071 153.9 318655 **304.6 1980 153.9 310 2108 153.2 322864 310 2015 153.2 320 2176 151.9 330631 320 2080 151.9 330 2244 150.8 338399 330 2145 150.8 340 2312 149.7 346165 340 2210 149.7 350 2380 148.7 353926 350 2275 148.7 * 353.2 2402 148.4 356399 353.2 2296 148.4 360 2448 148.7 364133 360 2340 148.7 370 2516 149.2 375483 370 2405 149.2 380 2584 149.7 386832 380 2470 149.7 390 2652 150.1 398180 390 2535 150.1 **399.0 2713 150.5 408393 "399.0 2594 150.5

Moment (inlb) 9009 18086 27187 36288 45388 54490 56862 65260 76324 87137 97967 108799 119632 130465 141297 152133 162981 173834 177503 181521 187588 193662 196623 202187 213040 223892 234745 245597 256449 267300 278151 289001 299852 304596 308620 316045 323469 330893 338311 340675 348068 358917 369765 380613 390376

CriticalfuelquantitiesformostforwardCGcondition (ncludes zero uel) Criticalfuelquantitiesformostaft CGcondition °

Weightsgivenare nominal weightsat 15' C (59 F) Forcalculation purposes. useWeightandMomentvalues.(CGsshownarederivedfromWeight andMomentvaluesandroundedtoonedecimalplace). 412-FMS-65-5-11-1

25


Table 5-11. Usablefuel - longitudinal- w/309.3 litre aux fuel In LH position(Metric) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (METRIC) Basic with LH 309.3 litre Aux Tank - 1510.4 litres Jet B, JP-4 (0.799 kgA) Jet A, A-1, JP-5, JP-8 (0.815 kgA) CG Quantity Weight Moment CG Weight Quantity (kg) (mm) (litres) (mm-kg) (mm) (kg) (litres) 32.0 3520 40 3520 114758 32.6 40 63.9 3534 80 3534 230430 80 65.2 95.9 3541 120 3541 346339 120 97.8 127.8 3545 462247 160 130.4 3545 160 159.8 3547 578156 200 163.0 3547 200 189.3 3548 *237.0 3548 685353 193.1 '237.0 191.8 3560 240 3560 696327 240 195.6 223.7 3680 280 3680 839843 280 228.2 255.7 3754 979007 320 260.8 3754 320 287.6 3807 1116853 360 293.4 3807 360 319.6 3849 400 3849 1254810 400 326.0 1392777 440 351.6 3884 358.6 3884 440 383.5 3913 480 1530743 3913 391.2 480 415.5 3937 520 1668711 423.8 3937 520 3959 447.4 560 3959 1806684 560 456.4 479.4 3977 600 3977 1944827 600 489.0 511.4 3994 2083059 640 521.6 3994 640 4000 656.3 524.4 2139449 4000 534.9 656.3 543.3 3943 680 3943 2185216 680 554.2 575.3 3856 2262497 720 586.8 3856 720 3778 607.2 2339857 760 619.4 3778 760 619.6 3750 775.5 2369900 3750 632.1 775.5 639.2 3765 2454459 800 652.0 3765 800 671.2 3787 2592684 840 684.6 3787 840 3808 703.1 880 2730905 717.2 3808 880 735.1 3827 920 3827 2869123 920 749.8 767.0 3844 960 3844 3007337 960 782.4 799.0 3860 1000 3145545 815.0 3860 1000 3874 1040 831.0 3283747 3874 847.6 1040 3888 862.9 1080 3888 3421946 1080 880.2 894.9 3900 1120 3560142 912.8 3900 1120 921.2 3907 "1153.0 3907 3671304 1153.0 939.7 926.8 3901 1160 3901 3687938 1160 945.4 3868 958.8 3782498 1200 978.0 3868 1200 3836 990.8 3877064 1240 1010.6 3836 1240 3807 1022.7 1280 3971622 1043.2 3807 1280 1054.7 3780 1320 4066111 3780 1075.8 1320 1068.2 3768 1337.0 4106161 3768 1089.6 1337.0 1086.6 3776 1360 4185772 1108.4 3776 1360 1118.6 3790 1400 3790 4323952 1400 1141.0 1150.6 3802 1440 4462119 3802 1173.6 1440 1182.5 3814 4600277 1480 1206.2 3814 1480 1206.8 3822 1510.4 4705202 3822 1231.0 1510.4

*

Moment (mm-kg) 112505 225907 339540 453172 566805 671898 682657 823355 959787 1094927 1230176 1365434 1500692 1635951 1771215 1906646 2042165 2097447 2142316 2218080 2293921 2323374 2406273 2541785 2677292 2812796 2948297 3083792 3219281 3354766 3490249 3599230 3615537 3708240 3800950 3893652 3986286 4025549 4103598 4239064 4374519 4509964 4612830

(Includes zero uel) CriticalfuelquantitiesformostforwardCGcondition Critcalluelquantitiestor mostaft CGcondition °

Weightsgivenare nominalweightsat 15 C (59°F) purposes,useWeightandMomentvalues.(CGsshownare derivedIromWeight Forcalculation andMomentvaluesandroundedtowholenumbers)

412-FMS-65-5-11-2

26


Table 5-12. Usable fuel - lateral - w/81.7 gal aux fuel in LH position (English) USABLE FUEL LOADING TABLE - LATERAL CG (ENGLISH) Basic with LH 81.7 US Gal Aux Tank - 399.0 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Quantity Weight CG Moment Quantity Weight CG (USGal) (lb) (inches) (in-lb) (USGal) (lb) (inches) 10 68 0.0 0 10 65 0.0 20 136 0.0 0 20 130 0.0 30 204 0.0 0 30 195 0.0 40 272 0.0 0 40 260 0.0 50 340 0.0 0 50 325 0.0 60 408 0.0 0 60 390 0.0 62.6 426 0.0 0 62.6 407 0.0 70 476 -0.2 -88 70 455 -0.2 80 544 -1.0 -565 80 520 -1.0 90 612 -2.1 -1262 90 585 -2.1 100 680 -2.8 -1928 100 650 -2.8 110 748 -3.5 -2587 110 715 -3.5 120 816 -4.0 -3245 120 780 -4.0 130 884 -4.4 -3903 130 845 -4.4 140 952 -4.8 -4559 140 910 -4.8 150 1020 -5.1 -5213 150 975 -5.1 160 1088 -5.4 -5855 160 1040 -5.4 170 1156 -5.6 -6491 170 1105 -5.6 173.4 1179 -5.7 -6706 173.4 1127 -5.7 180 1224 -5.8 -7042 180 1170 -5.8 190 1292 -5.5 -7170 190 1235 -5.5 200 1360 -5.4 -7305 200 1300 -5.4 204.9 1393 -5.3 -7369 204.9 1332 -5.3 210 1428 -5.4 -7695 210 1365 -5.4 220 1496 -5.6 -8331 220 1430 -5.6 230 1564 -5.7 -8966 230 1495 -5.7 240 1632 -5.9 -9602 240 1560 -5.9 250 1700 -6.0 -10237 250 1625 -6.0 260 1768 -6.1 -10872 260 1690 -6.1 270 1836 -6.3 -11507 270 1755 -6.3 280 1904 -6.4 -12142 280 1820 -6.4 290 1972 -6.5 -12777 290 1885 -6.5 300 2040 -6.6 -13411 300 1950 -6.6 304.6 2071 -6.6 -13661 *304.6 1980 -6.6 310 2108 -6.5 -13661 310 2015 -6.5 320 2176 -6.3 -13661 320 2080 -6.3 330 2244 -6.1 -13661 330 2145 -6.1 340 2312 -5.9 -13661 340 2210 -5.9 350 2380 -5.7 -13661 350 2275 -5.7 353.2 2402 -5.7 -13661 353.2 2296 -5.7 360 2448 -5.8 -14093 360 2340 -5.8 370 2516 -5.9 -14729 370 2405 -5.9 380 2584 -5.9 -15370 380 2470 -5.9 390 2652 -6.0 -16014 390 2535 -6.0 399.0 2713 -6.1 -16594 399.0 2594 -6.1

Moment (in-lb) 0 0 0 0 0 0 0 -85 -540 -1206 -1843 -2473 -3102 -3730 -4358 -4983 -5596 -6205 -6410 -6732 -6853 -6983 -7044 -7356 -7963 -8571 -9178 -9785 -10392 -10999 -11606 -12213 -12820 -13058 -13058 -13058 -13058 -13058 -13058 -13058 -13471 -14080 -14692 -15308 -15862

Criticalfuel quantityformostlateralCGcondition °

°

Weightsgivenarenominalweightsat 15 C (59 F) Forcalculation purposes,useWeightandMomentvalues,(CGsshownare derivedfromWeight andMomentvaluesandroundedtoonedecimalplace). 412-FMS-65-5-12-1

27


Usable fuel - lateral - w/309.3 litre aux fuel in LH position (Metric)

USABLE FUEL LOADING TABLE - LATERAL CG (METRIC) Basic with LH 309.3 litre Aux Tank - 1510.4 litres Jet A, A-1, J P-5, JP-8 (0.815 kg/l) Jet B, JP-4 (0.799 kg/l) Quantit y CG Moment Weight Quantity Weight CG (litres) (mm) (mm-kg) (ltres) (kg) (kg) (mm) 40 32.6 0 0 40 32.0 0 80 65.2 0 0 80 63.9 0 120 0 0 120 95.9 97.8 0 160 0 0 130.4 160 127.8 0 200 163.0 0 0 200 159.8 0 237.0 193.1 0 0 237.0 189.3 0 0 -24 240 195.6 240 191.8 0 280 228.2 -13 -2885 280 223.7 -13 320 -39 -10202 320 255.7 260.8 -39 360 293.4 -63 -18472 360 287.6 -63 400 326.0 -81 -26520 400 319.6 -81 440 -96 -34536 440 358.6 351.6 -96 480 -109 -42542 480 383.5 391.2 -109 520 423.8 -119 -50541 520 415.5 -119 560 456.4 -128 -58527 560 447.4 -128 600 489.0 -136 -66357 600 479.4 -136 640 521.6 -142 -74102 640 511.4 -142 656.3 534.9 -144 -77260 656.3 524.4 -144 680 554.2 -146 -81083 680 543.3 -146 720 586.8 -141 -82633 720 575.3 -141 760 619.4 -136 -84280 760 607.2 -136 775.5 -134 -84904 775.5 632.1 619.6 -134 800 652.0 -137 -89639 800 639.2 -137 840 684.6 -142 -97377 840 671.2 -142 880 717.2 -147 -105113 880 703.1 -147 920 749.8 -151 -112846 920 735.1 -151 960 782.4 -154 -120578 960 767.0 -154 1000 815.0 -157 -128309 1000 799.0 -157 1040 847.6 -161 -136042 1040 831.0 -161 1080 880.2 -163 -143771 1080 862.9 -163 1120 912.8 -166 -151498 1120 894.9 -166 * 1153. 939.7 -167 -157387 1153.0 921.2 -167 1160 945.4 -166 -157387 1160 926.8 -166 1200 978.0 -161 -157387 1200 958.8 -161 1240 1010.6 -156 -157387 1240 990.8 -156 1280 1043.2 -151 -157387 1280 1022.7 -151 1320 -146 -157387 1320 1075.8 1054.7 -146 1337.0 1089.6 -144 -157387 1337.0 1068.2 -144 1360 1108.4 -146 -161840 1360 1086.6 -146 1400 1141.0 -149 -169585 1400 1118.6 -149 1440 1173.6 -151 -177387 1440 1150.6 -151 1480 1206.2 -154 -185225 1480 1182.5 -154 1510.4 1231.0 -155 -191181 1510.4 1206.8 -155

Moment (mm-kg) 0 0 0 0 0 0 -24 -2829 -10002 -18109 -25999 -33858 -41707 -49549 -57378 -65054 -72647 -75743 -79491 -81011 -82625 -83237 -87879 -95466 -103049 -110631 -118210 -125790 -133371 -140949 -148524 -154297 -154297 -154297 -154297 -154297 -154297 -154297 -158663 -166256 -173905 -181589 -187427

CriticaluelquanttyformostlateralCG condition °

°

Weightsgivenare nominalweightsat 15 C. (59 F) Forcalculation purposes,useWeightandMomentvalues.(CGsshownare derivedfromWeight andMomentvaluesandroundedtowholenumbers). 412-FMS-65-5-12-2

28


Table 5-13. Usable fuel - longitudinal - w/81.7 gal aux fuel in RH position (English) USABLE FUEL LOADING TABLE - LONGITUDINAL CG (ENGLISH) Basic with RH 81.7 US Gal Aux Tank - 399.0 US Gal Jet A, A-1, JP-5, JP-8 (6.8 lb./Gal) Jet B, JP-4 (6.5 lb./Gal) Weight CG Quantity Weight CG Moment Quantity (US Gal) (inches) (USGal) (lb) (lb) (inches) (in-lb) 10 68 138.6 9424 10 65 138.6 20 136 139.1 18921 130 139.1 20 30 204 139.4 28442 195 30 139.4 260 40 272 139.6 37962 40 139.6 50 340 139.7 47483 50 325 139.7 60 408 139.7 57005 390 60 139.7 62.6 426 139.7 59486 407 139.7 '62.6 70 476 143.4 68272 143.4 70 455 80 544 146.8 79846 80 520 146.8 149.0 90 612 149.0 91159 90 585 100 680 150.7 102488 100 650 150.7 110 748 152.2 113821 715 110 152.2 120 816 153.4 125153 120 153.4 780 130 884 154.4 136486 130 845 154.4 140 952 155.3 147819 910 155.3 140 150 1020 156.0 159155 150 975 156.0 160 1088 156.7 170503 160 1040 156.7 170 1156 157.3 181857 157.3 170 1105 **173.4 1179 157.5 185696 ** 173.4 1127 157.5 180 1224 155.1 189899 180 1170 155.1 190 1292 151.9 196246 190 1235 151.9 200 1360 149.0 202600 1300 149.0 200 204.9 1393 147.7 205698 * 204.9 1332 147.7 210 1428 148.1 211519 1365 210 148.1 220 1496 149.0 222872 1430 149.0 220 230 1564 149.8 234226 230 1495 149.8 240 1632 150.5 245579 240 1560 150.5 250 1700 151.1 256932 1625 250 151.1 260 1768 151.7 268285 260 1690 151.7 270 1836 152.3 279637 270 1755 152.3 280 1904 152.8 290988 1820 152.8 280 290 1972 153.3 302340 1885 153.3 290 300 2040 153.8 313691 300 1950 153.8 304.6 2071 153.9 318655 **304.6 1980 153.9 310 2108 153.2 322864 310 2015 153.2 320 2176 151.9 330631 2080 151.9 320 330 2244 150.8 338399 2145 150.8 330 340 2312 149.7 346165 340 2210 149.7 148.7 350 2380 148.7 353926 350 2275 353.2 2402 148.4 356399 * 353.2 2296 148.4 360 2448 148.7 364133 2340 148.7 360 370 2516 149.2 375483 370 2405 149.2 380 2584 149.7 386832 380 2470 149.7 390 2652 150.1 398180 2535 150.1 390 **399.0 2713 150.5 408393 **399.0 2594 150.5

Moment (in-b) 9009 18086 27187 36288 45388 54490 56862 65260 76324 87137 97967 108799 119632 130465 141297 152133 162981 173834 177503 181521 187588 193662 196623 202187 213040 223892 234745 245597 256449 267300 278151 289001 299852 304596 308620 316045 323469 330893 338311 340675 348068 358917 369765 380613 390376

Critical fuel quantities or most forward CG condition (includes zero fuel) Critical fuel quantities for most aft CG condition

°

Weights given are nominal weights at 15 C (59' F) For calculation purposes, use Weight and Moment values, (CGs shown are derived trom Weight and Moment values and rounded to one decimal place)

412-FMS-65-5-13-1

29


Table 5-13. Usablefuel - longitudinal- w/309.3litre aux fuel in RH position(Metric) USABLE FUEL LOADING TABLE Basic with RH 309.3 litre Jet A, A-1, JP-5, JP-8 (0.815 kgA) Quantity Weight CG Moment (litres) (kg) (mm) (mm-kg) 40 32.6 3520 114758 80 65.2 3534 230430 120 97.8 3541 346339 160 130.4 3545 462247 200 163.0 3547 578156 *237.0 193.1 3548 685353 240 195.6 3560 696327 280 228.2 3680 839843 320 260.8 3754 979007 360 293.4 3807 1116853 400 326.0 3849 1254810 440 358.6 3884 1392777 480 391.2 3913 1530743 520 423.8 3937 1668711 560 456.4 3959 1806684 600 489.0 3977 1944827 640 521.6 3994 2083059 ** 656.3 534.9 4000 2139449 680 554.2 3943 2185216 720 586.8 3856 2262497 760 619.4 3778 2339857 * 775.5 632.1 3750 2369900 800 652.0 3765 2454459 840 684.6 3787 2592684 880 717.2 3808 2730905 920 749.8 3827 2869123 960 782.4 3844 3007337 1000 815.0 3860 3145545 1040 847.6 3874 3283747 1080 880.2 3888 3421946 1120 912.8 3900 3560142 **1153.0 939.7 3907 3671304 1160 945.4 3901 3687938 1200 978.0 3868 3782498 1240 1010.6 3836 3877064 1280 1043.2 3807 3971622 1320 1075.8 3780 4066111 *1337.0 1089.6 3768 4106161 1360 1108.4 3776 4185772 1400 1141.0 3790 4323952 1440 1173.6 3802 4462119 1480 1206.2 3814 4600277 1510.4 1231.0 3822 4705202

- LONGITUDINAL CG (METRIC) Aux Tank - 1510.4 litres Jet B, JP-4 (0.799 kg/1) Quantity Weight CG (litres) (kg) (mm) 40 32.0 3520 80 63.9 3534 120 95.9 3541 160 127.8 3545 200 159.8 3547 237.0 189.3 3548 240 191.8 3560 280 223.7 3680 320 255.7 3754 360 287.6 3807 400 319.6 3849 440 351.6 3884 480 383.5 3913 520 415.5 3937 560 447.4 3959 600 479.4 3977 640 511.4 3994 **656.3 524.4 4000 680 543.3 3943 720 575.3 3856 760 607.2 3778 *775.5 619.6 3750 800 639.2 3765 840 671.2 3787 880 703.1 3808 920 735.1 3827 960 767.0 3844 1000 799.0 3860 1040 831.0 3874 1080 862.9 3888 1120 894.9 3900 **1153.0 921.2 3907 1160 926.8 3901 1200 958.8 3868 1240 990.8 3836 1280 1022.7 3807 1320 1054.7 3780 *1337.0 1068.2 3768 1360 1086.6 3776 1400 1118.6 3790 1440 1150.6 3802 1480 1182.5 3814 1510.4 1206.8 3822

Moment (mm-kg) 112505 225907 339540 453172 566805 671898 682657 823355 959787 1094927 1230176 1365434 1500692 1635951 1771215 1906646 2042165 2097447 2142316 2218080 2293921 2323374 2406273 2541785 2677292 2812796 2948297 3083792 3219281 3354766 3490249 3599230 3615537 3708240 3800950 3893652 3986286 4025549 4103598 4239064 4374519 4509964 4612830

Criticalfuel quantitiesformostforwardCGcondition (includeszeo fuel) ** Criicalluel quantitiesformostaft CGcondition °

Weights givenare nominalweightsat 15 C.(59 F) Forcalculation purposes.useWeightandMomentvalues.(CGsshownare derivedfromWeight andMomentvaluesandroundedtoonedecimalplace) 412FMS-65-132

30


Table 5-14.

Usable fuel - lateral - w/81.7 gal aux fuel in RH position (English)

USABLEFUELLOADINGTABLE- LATERALCG (ENGLISH) Basicwith RH81.7 US Gal AuxTank- 399.0US Gal Jet A, A-1,JP-5,JP-8 (6.8lb./Gal) Jet B, JP-4 (6.5lb./Gal) Quantity Weight CG Moment Quantity Weight CG USGal) 10 20 30 40 50 60

(lb) 68 136 204 272 340 408

(inches) 0.0 0.0 0.0 0.0 0.0 0.0

62.6

426

70 80 90 100

476 544 612 680

110 120 130 140

Moment

(in- ) 0 0 0 0 0 0

(USGal) 10 20 30 40 50 60

(lb) 65 130 195 260 325 390

0.0

0

62.6

407

0.0

0

0.1 1.0 2.0 2.8

64 539 1235 1902

70 80 90 100

455 520 585 650

0.1 1.0 2.0 2.8

61 515 1181 1818

748 816 884 952

3.4 3.9 4.4 4.8

2561 3219 3876 4533

110 120 130 140

715 780 845 910

3.4 3.9 4.4 4.8

2448 3077 3705 4333

150 160

1020 1088

5.1 5.4

5186 5828

150 160

975 1040

5.1 5.4

4958 5571

170 173.4

1156 1179

5.6 5.7

6465 6680

170 173.4

1105 1127

5.6 5.7

6179 6385

180 190

1224 1292

5.2 4.8

6343 6216

180 190

1170 1235

5.2 4.8

6063 5942

200 204.9 210 220 230 240 250 260 270 280 290 300 *304.6 310 320 330 340 350 353.2 360 370 380 390 399.0

1360 1393 1428 1496 1564 1632 1700 1768 1836 1904 1972 2040 2071 2108 2176 2244 2312 2380 2402 2448 2516 2584 2652 2713

4.5 4.3 4.4 4.7 4.9 5.1 5.2 5.4 5.5 5.7 5.8 5.9 5.9 5.8 5.7 5.5 5.3 5.2 5.1 5.2 5.3 5.4 5.5 5.6

6080 6016 6342 6978 7613 8248 8883 9518 10154 10789 11423 12058 12307 12307 12307 12307 12307 12307 12307 12740 13374 14005 14635 15202

200 204.9 210 220 230 240 250 260 270 280 290 300 *304.6 310 320 330 340 350 353.2 360 370 380 390 399.0

1300 1332 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 1980 2015 2080 2145 2210 2275 2296 2340 2405 2470 2535 2594

4.5 4.3 4.4 4.7 4.9 5.1 5.2 5.4 5.5 5.7 5.8 5.9 5.9 5.8 5.7 5.5 5.3 5.2 5.1 5.2 5.3 5.4 5.5 5.6

5812 5751 6062 6670 7277 7884 8492 9099 9706 10313 10919 11526 11764 11764 11764 11764 11764 11764 11764 12178 12784 13388 13989 14531

(inches) 0.0 0.0 0.0 0.0 0.0 0.0

(in-.) 0 0 0 0 0 0

Critical fuelquantityfor mostlateral CGcondition

Weightsgivenare nominalweightsat 15 C. (59 F) Forcalculationpurposes,useWeightan Moment values,(CGs shown are derivedfromWeight andMomentvaluesandroundedto onedecimalplace). 412-FMS-65-5-14-1

31


Usable fuel - lateral - w/309.3 litre aux fuel in RH position (Metric)

USABLEFUELLOADINGTABLE- LATERALCG (METRIC) Basicwith RH309.3litre AuxTank - 1510.4litres Jet A, A-i, JP-5,JP-8 (0.815kgA) Jet B. JP-4 (0.799kg/) Quantity Weight CG Moment Quantity Weight CG

Moment

(lires) 40 80 120 160 200 237.0 240

(kg) 32.6 65.2 97.8 130.4 163.0 193.1 195.6

(mm) 0 0 0 0 0 0 0

(mm-kg) 0 0 0 0 0 0 -24

((Itres) 40 80 120 160 200 237.0 240

(kg) 32.0 63.9 95.9 127.8 159.8 189.3 191.8

(m) 0 0 0 0 0 0 0

280 320 360 400 440 480 520 560 600 640 656.3 680 720 760 775.5 800 840 880 920 960 1000

(mm-kg) 0 0 0 0 0 0 -24

228.2 260.8 293.4 326.0 358.6 391.2 423.8 456.4 489.0 521.6 534.9 554.2 586.8 619.4 632.1 652.0 684.6 717.2 749.8 782.4 815.0

11 38 62 80 95 108 119 128 135 141 144 132 122 113 110 114 119 125 130 134 138

2589 9898 18168 26216 34232 42239 50237 58223 66053 73798 76956 73133 71583 69937 69313 74048 81786 89522 97255 104986 112718

280 320 360 400 440 480 520 560 600 640 656.3 680 720 760 775.5 800 840 880 920 960 1000

223.7 255.7 287.6 319.6 351.6 383.5 415.5 447.4 479.4 511.4 524.4 543.3 575.3 607.2 619.6 639.2 671.2 703.1 735.1 767.0 799.0

11 38 62 80 95 108 119 128 135 141 144 132 122 113 110 114 119 125 130 134 138

2538 9704 17811 25701 33560 41409 49251 57080 64756 72349 75446 71698 70178 68564 67952 72594 80181 87764 95345 102925 110505

1040

847.6

142

120451

1040

831.0

142

118086

1080 1120 '1153.0 1160 1200 1240 1280

880.2 912.8 939.7 945.4 978.0 1010.6 1043.2

146 149 151 150 145 140 136

128180 135907 141796 141796 141796 141796 141796

1080 1120 '1153.0 1160 1200 1240 1280

862.9 894.9 921.2 926.8 958.8 990.8 1022.7

146 149 151 150 145 140 136

125663 133239 139012 139012 139012 139012 139012

1320

1075.8

132

141796

1320

1054.7

132

139012

1337.0 1360 1400 1440 1480 1510.4

1089.6 1108.4 1141.0 1173.6 1206.2 1231.0

130 132 135 138 140 142

141796 146249 153970 161656 169323 175145

1337.0 1360 1400 1440 1480 1510.4

1068.2 1086.6 1118.6 1150.6 1182.5 1206.8

130 132 135 138 140 142

139012 143378 150947 158482 165999 171707

Criticalfuel quantityformostlateralCGcondition

Weightsgivenare nominalweightsat 15'C.(59' F) Forcalculation purposes,useWeightandMomentvalues,(CGsshownare derivedfromWeight andMomentvaluesandroundedto onedecimalplace) 412-FMS5-5-14-2

32

T N I MODELR P N O R ROTORCRAFT I O T FLIGHT MANUAL F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-66.4

SUPPLEMENT

GLOBAL POSITIONING SYSTEM (TRIMBLE 2101 APPROACH) WITH ELECTRICAL/MECHANICAL HSI DISPLAY 412-899-378 CERTIFIED 19 JUNE 1998

This supplementshall be attachedto Model412EPFlight Manual (BHT-412-FM-4)when Trimble 2101 Approach Global Positioning System has been installed in accordancewith 412-899-378.

Information contained herein supplements information of basic Flight Manual.For Limitations, Procedures,and Performance Data not contained in this supplement, consult basic Flight Manual.

19 JUNE 1998


NOTICEPAGE


Manufacture's Data portion of this supplement Is proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron Inc.



LOG OF REVISIONS

Original ......................0...................19 JUN 98

LOG OF PAGES

REVISION NO.

PAGE

PAGE

FLIGHTMANUAL

Title - NP................................................ 0 A - B ....................................................... 0 I- ii ........................................................ 0 1- 2 ........................................................ 0 3/4............................................................. 0

REVISION NO. MANUFACTURER'S DATA

5 - 6 ....................................................... 0 7/8 ........................................................... 0

NOTE

Revisedtext is indicatedby a black vertical line. Insert latest revision pages; disposeof supersededpages. A



Original ....................... 0 ................... 19 JUN 98

APPROVED

MANAGER

ROTO CRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170

DATE


GENERAL INFORMATION

Global Positioning System is a global coverage satellite based navigation system. It provides standard positioning service (SPS)to all equipped users. Trimble2101Approach is a long range, GPS based RNAVsystem consisting of a receiver/display unit and an antenna. This system provides pilot with navigationalguidance and position information in three dimensions: latitude, longitude, and altitude. Every 28 days Trimble makes Jeppesen data base NavDatacard updates available to Trimble 2101ApproachGPSusers.

Trimble 2101 Approach has been designed to meet accuracy specification of VFR/IFRenroute oceanic and remote enroute domestic operation within U.S. National Airspace System, North Minimum NavigationPerformanceSpecification(MNPS)Airspaceand latitudes bounded Atlantic by 74° Northand 60° South using WGS-84(or NAD83) coordinatereferencedatum in accordance with criteria of AC 20-138,AC 91-49,and AC 120-33.Navigation data is based upon use of GPS operated by the United States. System accuracy is predicated on system receiving usable signals. NOTE Helicopters using GPSfor oceanic IFR operations may use Trimble 2101Approach to replace other meansof "long-range" navigation.A single Trimble 2101Approach GPSinstallation may also be used on short oceanic routes which require only one meansof "longrange" navigation.


ABBREVIATIONS

APR

Approach

CRS

Course

DTK

Desired Track

FAF

Final Approach Fix

GPS

Global Positioning System

HLD

Hold (DME)

IFR

Instrument Flight Rules

ILS

Instrument Landing System

LDA

Localizer Type Directional Aid

LOC/BC

Localizer Backcourse

MLS

Microwave Landing System

MSG

Message

OBS

Omni Bearing Selector

PTR

Pointer

RAIM

Receiver Autonomous Integrity Monitoring

SDF

Simplified Directional Facility

WPT

Waypoint

T N Section 1 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-66.4

FAA APPROVED

1-1.

INTRODUCTION

Limitations described in this supplement are in addition to those described in Basic Flight Manual and apply only to Model 412EP helicopters with GPS kit installed. Trimble 2101 Approach GPS Pilot's Guide, (Publication Number 81451) dated November 1996 shall be immediately available to flight crew whenever navigation is predicated on use of system.

Software number of Pilot's Guide shall match System Code annunciated on System Code page.

1-3.

TYPES OF OPERATION

Trimble 2101 Approach GPS Navigator is for single pilot IFR Enroute approved Navigation only. IFR GPS Approaches are prohibited.

System shall utilize Software number 0237.

GPS control box shall be located within pilot reach and view, on pilot side of center console, in approximately center position.

IFR En Route navigation is prohibited unless pilot verifies currency of data base or verifies each selected waypoint for accuracy by reference to current approved data.

GPS equipment for IFR navigation is for use as a supplemental navigation system.

1-6. WEIGHT OF GRAVITY 1-6-B.

AND

CENTER

CENTEROF GRAVITY


T N Section2 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

2-4. INTERIOR START CHECK

BHT-412-FMS-66.4

AND

PRE-

2-4-A-1. LOWERPEDESTAL

FORCE TRIM switch - ON, cover down.

GPS POWER switch - OFF.

2-6.

2-6-H.

SYSTEMS CHECK GPS CHECK

GPS POWER switch - ON.

Verify GPS annunciators (APR, HLD, WPT and MSG) are illuminated.

After Self-Test is passed, verify and adjust if needed, correct barometer setting is displayed on the "Air Data" page; to do so, follow procedure described in Trimble 2101 Approach Pilot Guide.

Verify proper barometric altitude is displayed on Altitude page; to do so, follow procedure described in Trimble 2101 Approach GPS Pilot's Guide. This altitude should be within 100 ft. of actual geographic altitude of A/C.

2-9.

IN-FLIGHTOPERATIONS

2-9-A SELECTIONOF GPSTO HSI CRS PTR switch annunciator GPS legend is illuminated.

-

Cycle until

Normal operating procedures are outlined in Trimble 2101 Approach GPS Pilot's Guide, (Publication Number 81451) dated November 1996.

2-9-B.

SYSTEMANNUNCIATIONS

2-9-B-1. APPROACH(APR) HOLD(HLD)/ WAYPOINT(WPT) MESSAGE(MSG)/

HLD (white) annunciator on Pilot and Copilot Instrument Panels (refer to TRIMBLE 2101 APPROACH Pilot Guide). WPT (amber) annunciator Pilot and Copilot Instrument Panels (refer to TRIMBLE 2101 APPROACH Pilot Guide). MSG (amber) annunciator on Pilot and Copilot Instrument Panels will flash to alert pilot(s) of (refer to requiring attention a situation TRIMBLE 2101 APPROACH Pilot Guide).

T N * Section 3 I R P N O R I O T F C U T O OD N R E P E E R Section 4 SCR N G O N I R W O E I F V NLY O BHT-412-FMS-66.4

FAA APPROVED

3-15. GPS MESSAGES RAIMrelated messages:

When one or more following RAIM related messagesappears in flight: INDICATIONS:

"RAIM PREDICTEDN/A"

"APPROACHRAIMLOST" "RAIM ERROR"

* PROCEDURE:

IFR Enroute - Continue to navigate using TRIMBLE2101 APPROACH,while verifying position using another IFR approved navigation system, or revert to alternate meansof navigation. Other messages:

Refer to Trimble 2101 APPROACH Pilot Guide. Discontinue IFR GPS operations, If message indicates GPS malfunction or degradationof capability.


3/4

T Section5 N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUI FACTURER'S DATA

BHT-412-FMS-66.4


5

T N Section1 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-66.4

DATA MANUFACTURER'S

1-17. NAVIGATION SYSTEM AND INSTRUMENTS

1-17-B. OPTIONAL NAVIGATION EQUIPMENT

Annunciator

will illuminate to alert pilot of

TRIMBLE2101APPROACHon activationof: Approach mode (APR,green).

Holding mode of active waypoint (HLD,white).

Turn anticipation (WPT,amber).

1-17-B-9.GLOBAL POSITIONING SYSTEM (GPS) Remoteswitches:

OBSknob - Pilots HSIcourse select knob.

There is no provision for HSI analog course input in TRIMBLE 2101 APPROACH. To properly use deviation bar on HSI, pilot should manually set HSI pointer to course in TRIMBLE 2101 value prescribed APPROACHdisplay (DTK), after selecting GPS as navigation source with CRS PTR push-buttonannunciator.

APR/HLD/WPT/MSG - Approach, Hold, Waypoint and Message annunciator located on instrumentpanel, Figure 1-1.

Message(MSGamber).

NOTE

These four (4) lights will illuminate during system self-test.

CRSPTR- Navigation Source Switch/Annunciator located on instrument panel.

Use to select navigation source display on pilots HSI.Pressto select desired source.


BHT-412-FMS-66.4

COPILOT INSTRUMENT PANEL

412FS66.4-1-1

Figure 1-1 TRIMBLE 2101 APPROACH annunciator switch panels

7/8

T N I R P N O R I ROTORCRAFT O T FLIGHTC MANUAL F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-67.1 & 67.2

MODEL

SUPPLEMENT

PT6T-3BF ENGINE (30 MINUTE OEI RATING) 412-706-054

S/N 33001 - 33107 33108 -33213 AND 36001 -36019 CERTIFIED 18 DECEMBER 1998

This supplement shall be attached to Model 412 Flight Manual (BHT-412-FM-1 , BHT-412-FM-2, or BHT-412-FMS19.1) when PT6T-3BF Engine, 30 minute OEI rating kit is installed. Information contained herein supplements information of basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, or other applicable supplements; consult basic Flight Manual.

COPYRIGHT NOTICE 1998 COPYRIGHT BELL HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON INC. A DIVISION OF TEXTRON CANADA LTD.

ALLRIGHTS RESERVED

Bell Hellicopter A Subsidiary of Textron Inc. POST OFFICE BOX 482 · FORT WORTH,TEXAS 76101

18 DECEMBER MBER1998 1998


NOTICEPAGE


NP


LOG OF REVISIONS

Original

..........

DEC98 0....................18

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP................................................. 0 A-

PAGE

REVISION NO.

i/ii ............................................................... 0 1 - 8 ..........................................................0 9/10 ............................................................ 0

B ........................................................0

NOTE Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. A


FAA APPROVED


Original ........................0................... 18 DEC 98

APPROVED

MANAGER


DATE

DEC1 8 1998


FAA APPROVED

GENERAL INFORMATION

The PT6T-3B engine, in accordance with 412706-054, may be redesignated as PT6T-3BF to offer an increased 30 minute OEI rating. This 30 minute OEI rating represents improved OEI

capabilities while maximum continuous OEI performance is reduced from that presented in the basic Flight Manual or applicable supplement.

i/ii

T N I Section R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-67.1& 67.2

TWIN ENGINEOPERATION

INTRODUCTION

PT6T-3BF engine offers minute OEI rating.

an increased

30

WEIGHT

Actual weight changes shall be determined after kit is installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits.

POWER PLANT


ONE ENGINE INOPERATIVE (OEI) Continuous OEI

101.8%

30 minute OEI range

101.8 to 103.4%

Maximum OEI (30 minute)

103.4%

INTERTURBINE TEMPERATURE

TWIN ENGINEOPERATION

Pratt and Whitney Aircraft of Canada, Ltd. PT6T-3BF. NOTE


Maximum continuous

765°C

5 minute range

765 to 810° C

Maximum

810° C

Maximum start (2 seconds maximim above 960°C)

1090° C

ONE ENGINE INOPERATIVE (OEI)

Anytime an engine is operated in an OEI range, an entry shall be made in the helicopter logbook detailing the extent of operation in excess of twin engine takeoff power limits. This does not apply to approved ITT limits for starting or for power assurance checks.

GAS PRODUCER RPM (N1)

Maximum continuous OEI

810°C

30 minute OEI range

810 to 850° C

Maximum OEI

850°C

ENGINE OIL PRESSURE


NOTE

Gas producer indicator 212-075-037113 (or equivalent) must be installed prior to or concurrent with kit.

1


FAA APPROVED

ENGINEOIL TEMPERATURE Minimum

0° C

Continuousoperation 0 to 115° C Maximumfor MIL-L-7808 115°C oil Maximumfor MIL-L-23699 120°C oil Maximumfor DOD-L-85734 120°C oil

COMBININGGEARBOXOIL PRESSURE No changefrom basic manual.

COMBININGGEARBOXOIL TEMPERATURE

Minimum 0° C Continuousoperation 0 to 115°C Maximumfor MIL-L-7808 115°C oil Maximumfor MIL-L-23699 120°C oil Maximumfor DOD-L-85734120°C oil

INSTRUMENT MARKINGS AND PLACARDS

Refer to figure 1-1 for instrument range markings and figure 1-2 for placards and decals.

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I A R W O E I F V NLY O BHT-412-FMS-67.1

& 67.2


& 67.2

Instrument

markings

(Sheet

2 of 2)


FAA APPROVED

TWIN & OEI 101.8% 30 MIN OEI 103.4%

LOCATION: INSTRUMENT PANEL

412FS67.1-1-2

Figure 1-2

Placards and decals 5

T N Section 2 I R P N O R I O T F C U T O OD N R E P E E R Section 3 SCR N G O N I R W O E I F V NLY O FAAAPPROVED

BHT-412-FMS-67.1 & 67.2



T N Section 4 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O FAA APPROVED

BHT-412-FMS-67.1& 67.2

INTRODUCTION

Helicopter gross weight pounds.

Performance data presented herein are derived from engine manufacturer's specification power for PT6T-3BF engine less Installation losses.

Pressure altitude - 5,500 feet.

CLIMB AND DESCENT

Refer to figure 4-1 for increased single engine rate of climb - 30 minute power for helicopters with maximum gross weight of 11,600 pounds (refer to BHT-412-FM-1). Refer to figure 4-2 for single engine rate of climb - 30 minute power for helicopters with maximum gross weight of 11,900 pounds (refer to BHT-412-FM-2 and BHT-412-FMS19.1).

OAT-

10,500

0°C.

EXAMPLE: 1.

Enter rate of climb chart at 5,500 feet pressure altitude.

2.

Move right, horizontally, to intersect 0° OAT line.

3.

Descend vertically to intersect MAX GW line in lower portion of chart.

4.

Follow curvature of trend lines to intersect actual helicopter gross weight line of 10,500 pounds.

5.

Descend vertically to bottom of chart and read 500 feet per minute rate of climb.

PROBLEM:

What is maximum rate of climb for following conditions?

7


BHT-412-FMS-67.1 & 67.2

SINGLEENGINERATEOF CLIMB

30 MINUTEPOWER ENGINERPM97% GENERATOR 150 AMPS

70 KIAS HEATEROFF

WITHALLDOORSOPENOR REMOVED:

1. CLIMBSPEEDIS 60KIAS 2. RATEOFCLIMBWILL DECREASE 275FT/MIN

20

-6

18 -

5

-4

u

o

-3

I

-

2

0 12

MAXGW ='1600

.

52

11

-48

10

44

98

7

-2

0

2

4

-1

0

1

2

6 8 10 12 14 16 RATE OF CLIMB - 100 FT/MIN 3 4 5 6 7 RATE OF CLIMB - M/S

8

18

9

20

10

22

32

11

412FS67

Figure 4-1. Single engine rate of climb - 30 minute power (11,600 pounds) (Sheet 1 of 2)

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O , i i

-2

0

2

4

6 10 12 14 16 18 20 RATEoF8Cums - ‘oil FTlMlN

22

T N I R P N O R ROTORCRAFT I O T MANUAL FLIGHT F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-68.3

SUPPLEMENT

PT6T-3BG ENGINE(30 MINUTE OEI RATING) 412-706-054

S/N 36020 -

36080 *

AND HELICOPTERS MODIFIED PER 412-570-001-103 CERTIFIED 18 DECEMBER 1998 This supplement shall be attached to Model 412 Flight Manual (BHT-412-FM-3or BHT-412-FMS-34.2)when PT6T-3BGEngine, 30 minute OEI rating kit has been installed. *HelicoptersS/N36081through 36086are equipped with PT6T-3Dengines.

Information contained herein supplements information of basic Flight Manual. For Limitations,Procedures,and PerformanceData not contained in this supplement, or other applicable supplements, consult basic Flight Manual.

Bell

COPYRIGHT NOTICE 1998 COPYRIGHT BELL HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON INC. A DIVISION OF TEXTRON CANADA LTD. ALL RIGHTS RESERVED

Helicopter A Subsidiary of Textron

POST

OFFICE BOX 482

FORT WORTH,TEXAS 76101

18 DECEMBER1998

Inc.


NOTICEPAGE

Additional copies of this publication maybe obtained by contacting: CommercialPublication Distribution Center Bell HelicopterTextronInc. P.O. Box 482 Fort Worth, Texas76101-0482

NP


LOG OF REVISIONS

Original .................... 0.................... 18 DEC 98

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title - NP .................................................

0

PAGE

REVISION NO.

A - B .........................................................0 ii ............................................................... 0 1 - 8 ..........................................................0

NOTE



BHT-412-FMS-68.3


Original ........................0 ................... 18 DEC 98

APPROVED

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AIATION ADMINISTRATION FT. WORTH, TX 76193-0170

DATE DEC 18

1998


FAA APPROVED

GENERAL INFORMATION

The PT6T-3BEengine, in accordance with 412-706-054,may be redesignated as PT6T3BG to offer an increased 30 minute OEI rating. This 30 minute OEI rating represents

improved OEI capabilities while maximum continuous OEI performanceis reduced from that presented in the basic Flight Manual or applicablesupplement.

i/ii


FAA APPROVED

1-1.

INTRODUCTION

PT6T-3BG engine offers minute OEI rating.

an increased

1-13-A-1. TWIN ENGINE OPERATION

30

1-6. WEIGHT

No change from basic manual. 1-13-A-2. (OEI)

ONE

ENGINE

INOPERATIVE

Continuous OEI

101.8%

Actual weight changes shall be determined after kit is installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits.

30 minute OEI range

101.8 to 103.4%

Maximum OEI (30 minute)

103.4%

1-13. POWER PLANT


Pratt and Whitney PT6T-3BG

Aircraft of Canada, Ltd.

1-13-C-1. TWIN ENGINE OPERATION

NOTE


Anytime an engine is operated in an OEI range, an entry shall be made in the helicopter logbook detailing the extent of operation in excess of twin engine takeoff power limits. This does not apply to approved ITT limits for starting or for power assurance checks.

1-13-A. GAS PRODUCERRPM(N1) NOTE

Gas producer indicator 212-075-037113 (or equivalent) must be installed prior to or concurrent with kit.

Maximum Continuous

765°C

5 minute range

765 to 810° C

Maximum

810° C

Maximum start (2 seconds maximum above 960°C)

1090°C

1-13-C-2. ONE ENGINE INOPERATIVE (OEI) Maximum Continuous OEI

810°C

30 minute OEI range

810 to 850°C

Maximum OEI

850°C

1-13-F. ENGINE OIL PRESSURE No change from basic manual.

1


BHT-412-FMS-68.3

1-13-J. COMBININGGEARBOX OIL TEMPERATURE

1-13-G. ENGINEOIL TEMPERATURE Minimum


0°C 0 to 115°C 115°C

Maximumfor MIL-L-7808 oil Maximumfor MIL-L-23699 120°C oil Maximumfor DOD-L-85734 120°C oil

Minimum

0°C

0 to 115°C Continuous operation Maximumfor MIL-L-7808 115°C oil Maximumfor MIL-L-23699 120°C oil Maximumfor DOD-L-85734 120°C oil

1-13-H. COMBININGGEARBOX OIL PRESSURE


No changefrom basic manual.

Refer to figure 1-1 for instrument range markings and figure 1-2 for placards and decals.


BHT-412-FMS-68.3

Figure

l-l

Instrument

markings

(Sheet

1 of 2)


412FS68.3-1

Figure

l-l

Instrument

markings

(Sheet

2 of 2)


BHT-412-FMS-68.3

TWIN & OEI 101.8% 30 MIN OEI 103.4%

LOCATION: INSTRUMENT PANEL

412FS68.3-1-2

Figure 1-2

Placards and decals

5

T N Section 2 I R P N O R 0 I O T F C U T O OD N R E P E E R Section 3 SCR N G O N I R W O E I F V NLY O FAAAPPROVED

BHT-412-FMS-68.3




FAA APPROVED

Helicopter gross weight pounds.

4-1. INTRODUCTION

Pressure altitude - 5,500 feet.

data presented herein are Performance from engine manufacturer's derived specification power for PT6T-3BG engine less installation losses.

4-7.

CLIMB AND DESCENT

Refer to figure 4-1 for increased single engine rate of climb, 30 minute power.

OAT - 0 ° C

EXAMPLE: 1.

Enter rate of climb chart at 5,500 feet pressure altitude.

2.

Move right, horizontally, to intersect 0 ° OAT line.

3.

Descend vertically to intersect MAX GW line in lower portion of chart.

4.

Follow curvature of trend lines to intersect actual helicopter gross weight line of 10,500 pounds.

5.

Descend vertically to bottom of chart and read 500 feet per minute rate of climb.

PROBLEM:

What is maximum rate of climb following conditions?

10,500

for

7


BHT-412-FMS-68.3

SINGLEENGINERATEOF CLIMB

70 KIAS HEATEROFF

30 MINUTEPOWER ENGINERPM97% 150AMPS GENERATOR

WITHALL DOORSOPENORREMOVED:

1. CLIMBSPEEDIS 60KIAS 275FT/MIN 2. RATEOFCLIMBWILL DECREASE

20

-6

18

-

-5 -

-16

14

U

-

RATE OF CLIMB - M/S

Figure 4-1. Single engine rate of climb - 30 minute power

412FS683- 4-1

T N I R P N O R I ROTORCRAFT O T FLIGHTC MANUAL F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-69.4

MODEL

SUPPLEMENT

DUAL DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM, SEARCH AND RESCUE (SAR) 412-705-008

S/N 36218AND SUB CERTIFIED 29 JANUARY 1999

This supplementshall be attachedto Model 412EPFlight Manual when Dual Digital Automatic Flight Control System, Search and Rescue (SAR) modification Is installed.

Information contained herein supplementsinformation of Basic Flight Manual. For Limitations, Procedures, and PerformanceDatanot containedin this supplement, consult Basic Flight Manual.

29 JANUARY 1999


NOTICEPAGE




NP


LOG OF REVISIONS

Original............................

. 29 JAN99

LOG OF PAGES

PAGE

REVISION NO.

T itle............................................................ 0 NP.............................................................. 0 A - B ........................................................0 i- ii .......................................................... 0 1 - 2 .........................................................0 3/4 ..............................................................0

PAGE

REVISION NO.

5 - 1 0 ........................................................0 11/12 ..........................................................0 13 - 20 ......................................................0 MANUFACTURER'S DATA 21 - 24 ......................................................0

NOTE Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. A



0 ....................29JAN 99 Original ........................

APPROVED

MANAGER

OFFICE CERTIFICATION ROTORCRAFT FEDERALAVIATIONADMINISTRATION TX 76193-0170 FT.WORTH, B

DATE


GENERAL INFORMATION

The 412 SAR configuration includes the Honeywell SPZ-7600 Integrated Flight Control System (DFZ-760 Digital AFCS, EDZ-705 EFIS, Primus 700 Weather Radar, and AA-300 Radio Altimeters), Flight Management System (Universal Avionics UNS-1C/D), Doppler (Racal D91 or CMC 2012C). Refer to appropriate manufacturer's publication for operations.

ABBREVIATIONS,ACRONYMS AND PLACARDING

EADI

-

EFIS Attitude Director Indicator

EEPROM

-

Electrical Erasable Programmable Read Only Memory

EFIS

-

Electronic Flight Instrument System

EHSI

-

EFIS Horizontal Situation Indicator

EMER

-

Emergency

ENC

-

Encoding

ADF

-

Automatic Direction Finder

EXT

-

Extinguisher

AGL

-

Above Ground Level

FD

-

Flight Director

AP

-

Autopilot

FDSEL

-

Flight Director Select

APR

-

Approach

FMS

-

Flight Management System

BLWR

-

Blower

fpm

-

Feet Per Minute

CDI

-

Course Deviation Indicator

FWD

-

Forward

CDU

-

Control Display Unit

GPS

-


CLTVYAW

-

Collective-yaw

HDG

-

Heading

HOV

-

Hover

COMM

-

Communication

HSI

-

Horizontal Situation Indicator

CPL

-

Couple

HTR

-

Heater

CPLT

-

Copilot

IMC

-

CPU

-

Central Processor Unit

Instrument Meteorological Conditions

CRS

-

Course

IAS

-

Indicated Airspeed

-

Intercommunication System

DCPL

-

Decouple

ICS

DECEL

-

Deceleration

IDB

-

Internal Data Base

DEG

-

Degrees

IDENT

-

Identification

-

Instrument Landing System

DH

-

Decision Height

ILS

DME

-

Distance Measuring Equipment

LE

-

Level

LF

-

Low Frequency

DN

-

Down

LH

-

Left Hand

DPLR

-

Doppler

LNAV

-

Long Range Navigation

DR

-

Dead Reckoning


MKR BCN

-

Marker Beacon

STBY

-

Standby

MOT

-

Mark On Target

SX

-

Selected Cross-track

MSG

-

Message

SYNC

-

Synchronize

TACAN

-

NAV

-

Navigation

Tactical Air Navigation

TAS

-

NDB

-

Nondirectional Beacon

True Airspeed

NM

-

TORQ/ TRQ

-

Torque

Nautical Mile

PRI

PWR

-

Primary

TRANS

-

Transfer

-

Power

VEL HLD

-

Velocity Hold

RAD ALT

-

Radar Altimeter

VLF

-

Very Low Frequency

REL

-

Release

VOR

-

VHF Omnidirectional Range

RH

-

Right Hand

WP/WPT

-

Waypoint

RNAV

-

Radio Navigation

WSHLD

-

Windsheild

SAR

-

Search and Rescue

WX

-

Weather

SAR Modes

-

Modes of operation (MOT, APR 1 and 2, CLIMB)

XPDR

-

Transponder

SBY ATT

-

Standby Attitude

SG REV

-

Symbol Generator Reversionary


1-1.

INTRODUCTION

Limitations specified in Basic Flight Manual, BHT-412-FMS-38.4, and other supplements remain applicable. Operation of SPZ-7600 system does not relieve flight crew of responsibility to monitor and comply with all helicopter limitations.

BHT-412-FMS-69.4

In addition to basic equipment required for certification and EFIS with four axis AFCS (412-705-009/024), the following items shall be operational for SAR operations (IFR): Collective adjustable friction or collective force trim Two operational EFIS systems

Two operational autopilot (AP) systems

1-3.

TYPES OF OPERATION

Overwater (SAR Mode) operations are approved for Day/Night, VFR/IFR conditions when conducted in accordance with appropriate operating rules and required equipment.

Two operational radar altimeter systems Approved, operational Flight Management System (FMS)

Approved, operational doppler system.

NOTE

IFR operations are not approved with:

External cargo loads using SAR modes Hoist operations (unless using SAR modes) Any doors open (unless using SAR modes).

1-4.

FLIGHT CREW

For overwater IFR automatic approach and hover operations, two pilots are required.

For all SAR operations with hoist installed, a hoist operator is required.

1-5.

1-5-A.

CONFIGURATION

REQUIRED EQUIPMENTOVERWATERSAR MODE OPERATIONS(IMC)

When referring to required equipment lists in Basic Flight Manual, note that EFIS replaces ADIHSI.

After initiation of SAR mode, if partial equipment failure occurs, pilot in command may determine minimum equipment required to complete mission using automatic approach and hover system.

Regardless of visual limitations, the following items shall be available for overwater operations: Approved life vests for all occupants Life raft

Weather radar with ground mapping capability.

1


BHT-412-FMS-69.4

1-5-C.

OPTIONAL EQUIPMENT

CAUTION

Emergency floats

IF ONE AUTOPILOT FAILS, SECOND AUTOPILOT WILL REMAIN COUPLED HOWEVER, MODES. IN SAR SUBSEQUENT ELECTRICAL FAILURES OR AFCS FAILURES MAY RESULT IN TOTAL LOSS OF AIRCRAFT STABILIZATION AND FLIGHT DIRECTOR GUIDANCE.

Hoist

Navigation graphics computer (DATANAV) Nightsun FLIR

Maximum Continuous Power kit NOTE

rules operating Appropriate equipment optional require installed.

may be

Pilot shall have hands lightly on controls when coupled to SAR modes below 100 feet AGL, dual autopilot, or anytime when coupled single autopilot.

CAUTION

1-6.

WEIGHT AND CENTER OF GRAVITY

Actual weight change shall be determined after optional kits are installed and ballast readjusted, if necessary, to return empty weight CG to within allowable limits. SAR operations shall be limited to OGE hover weights.

1-23. FLIGHT OPERATIONS 1-23-A.

SAR MODES,VFR / IFR

Approach 2 operations into the wind.

shall be conducted

SAR modes may be selected anytime within engagement limits.

Minimum airspeed during Approach 1 and MOT prior to Approach 2 or final MOT descent shall be 60 KIAS.

DURING NORMAL AFCS COUPLED OPERATION, SYSTEM TOLERANCES MAY ALLOW BASIC HELICOPTER LIMITATIONS TO BE EXCEEDED FOR SHORT DURATIONS. PILOT REMAINS RESPONSIBLE TO MONITOR AND CORRECT FOR ANY EXCEEDANCES.

When collective is coupled, the FZ-706 Flight Control Computer limits engine ITT or MAST or TORQUE, based on engine series transmission configuration installed. CAUTION

FLIGHT CONTROL THE FZ-706 COMPUTER DOES NOT RECOGNIZE ENGINE FAILURE. FOLLOWING ENGINE FAILURE, COMPUTER WILL CONTINUE TO LIMIT REMAINING ENGINE TO APPROPRIATE LIMITS DISENGAGES PILOT UNTIL OR COUPLING COLLECTIVE MANUALLY OVERRIDES COLLECTIVE LIMITING. UNDER CERTAIN LOW OAT CONDITIONS, THE AFCS COMPUTER MAY EXCEED SINGLE ENGINE TORQUE LIMIT.


BHT-412-FMS-69.4

1-24. FLIGHT MANAGEMENT SYSTEM (FMS) Refer

to

appropriate

Flight

Manual

Supplementfor installed Flight Management Systemlimitations and locations. Prior to flight, verify by CPU readout on initialization page (of installed FMS) that

approved computer program (or subsequently approved program) has been installed and applicable operator's manual is available.

1-26. WEATHER RADAR

Referto WeatherRadarOperator'sManualfor weather radarsystem limitations.

3/4


2-1.

INTRODUCTION

Normal procedures specified in BHT-412-FM-4 and BHT-412-FMS-38.4are applicable.

Referto Section 1 of this and other applicable

supplements, and the Basic Flight Manual, for subsystems restrictions.

Refer to Weather Radar Operator's Manual for weather radar system normal procedures (if installed).

2-2.

2-2-A.

2-3.

2-3-A.

TAKEOFFAND LANDING DATA

WEIGHTAND BALANCE

Determine proper weight and balance of helicopter for optional equipment installed.

Consult Manufacturer'sDatafor instructions. WARNING

HOIST LOAD (IF HOIST INSTALLED) CAN CAUSE LONGITUDINAL OR LATERAL CG TO EXCEED LIMITS. GROSS WEIGHT AND CENTER OF GRAVITY SHALL BE COMPUTED TO ASSURE LOADING WITHIN APPROVED LIMITS.

PREFLIGHT CHECK EXTERIORCHECK

Doppler housing - Condition.

Refer to appropriate Flight Manual Supplement for installed Flight Management System (FMS) exterior checks.

2-4.

FLIGHT PLANNING

SAR operations are restricted to OGE hover weights.

2-2-B.

BHT-412-FMS-69.4

INTERIOR AND PRESTART CHECK

Life vests and life raft - Installed serviceable. (Overwater operations) ICS - Check communication and hoist operator.

2-6.

2-6-G.

and

between pilot

SYSTEMS CHECK

ELECTRICALSYSTEMS CHECK

FMS power -

ON.

Verify approved, current program / data base appears on initialization page. Refer to appropriate Flight Manual Supplement for installed Flight Management System (FMS). Doppler power - ON.

Doppler modes are selected by FMS.

SELF TEST - Mode is inhibited coupled to autopilot.

when

Mode/s - Selected.

5


BHT-412-FMS-69.4

Refer to FMS Operators Manual for selection procedures. EHSI course display - Set to LNAV.

2-6-H.

2-9.

HYDRAULIC SYSTEM CHECK NOTE

Cyclic and collective frictions should be set to minimum for 4 axis mode operations.

2-6-J.

If error codes are present, press SBY (or cyclic NAV STBY switch) once to view each code until END 1 is displayed (refer to Section 3 for list of error codes).

BEFORE TAKEOFF

Set as

Manual Flight Refer to appropriate Supplement for installed Flight Management System (FMS).

Doppler - After 30 seconds, check operation by pressing HOV switch on EFIS (DC-811) display controller and verify doppler is valid. NOTE

Amber Doppler display and no velocity bars - use light lateral cyclic inputs to gently rock helicopter. Velocity bars 6

2-9-D.

SAR OPERATIONS

2-9-D-1.

APPROACH TO HOVER (APR 1)

FORCE TRIM switch - ON. CLTC/YAWTRIM switch -

ON.

AP2 switch - ON.

Press SBY to continue with Level 2 test. Upon completion of Level 2 test and AL-300 reads END 2, press TEST switch to exit AFCS test.

Flight Management System (FMS) desired.

IN-FLIGHTOPERATIONS

AP1 switch - ON.

AFCS CHECK

Upon completion of Level 1 test, verify AP1 and AP2 lights extinguish and AL-300 reads END 1.

2-7.

doppler and amber displayed extinguished validates doppler.

APR1 switch illuminates.

-

Press.

Verify

light

on EADI and Verify APR1 annunciation monitor 500 fpm rate of descent. NOTE

engage will mode ALT RAD automatically then automatically, disengage when descent is initiated.

If selected from 200 feet to 2000 feet AGL, helicopter will decelerate to 60 KIAS and descend to 200 feet AGL.

Monitor automatic level off at 200 feet AGL, according to radio altimeter, and airspeed hold at 60 KIAS. Confirm RAD ALT and IAS annunciation ON. Confirm EADI and APR1 annunciation OFF. NOTE

During approach profile, airspeed may be captured at anytime by momentarily pressing cyclic FORCE TRIM release switch or longitudinal beep switch. RAD ALT may be captured anytime by momentarily pressing CLTV-YAW TRIM


BHT-412-FMS-69.4

REL switch or beeping CLTV-YAW TRIM switch UP or DN.

pressing cyclic FORCE TRIM release switch or longitudinal beep switch.

During APR1 engagement, heading can be changed by engaging HDG switch on mode selector and selecting desired heading with heading control, or by using left/right cyclic beep to change heading manually.

During approach, roll axis will maintain a zero lateral doppler velocity. Reference can be changed using lateral beep switch. Heading can be changed using CLTV-YAW TRIM switch in L or R position, and will produce a flat uncoordinated turn to selected heading.

APR1 mode will be cancelled by:

Completion of APR1 procedure

RAD ALT displays - Monitor capture of selected altitude, primarily EHSI vertical display.

Pressing APR button on SAR engage panel

Selecting SBY through either MS-700 Flight Director Mode Selector or remote NAV STBY button on either cyclic

Doppler groundspeed hover display.

Selecting MOT mode.

EADI - Ensure VEL HLD and RAD ALT annunciators are ON and APR2 is OFF when approach 2 profile is complete.

2-9-D-2.

APPROACH TO HOVER(APR 2)

APR2 can be engaged at altitudes between 20 feet AGL and 250 feet AGL, with airspeed less than 65 KIAS. APR2 will not engage if these conditions are not met.

Press. Verify light illuminates.

HOV switch (EFIS display controller) Monitor doppler ground velocity.

Monitor on EHSI

APR2 mode will be cancelled by:

NOTE

APR switch -

-

ON.

AL-300 - Note hover altitude displayed. If desired, select new hover altitude between 50 and 150 feet AGL. Descent - Monitor and check APR2 and IAS annunciation on EADI. At 175 feet AGL, IAS hold will disengage and a slow deceleration to a hover will commence. NOTE

During approach, VEL HLD may be engaged anytime by momentarily

Completion of APR2 procedure.

Pressing APR button on SAR engage panel.

Selecting SBY through either MS-700 Flight Director Mode selector or remote NAV STBY button on either cyclic. Selecting MOT mode.

Selecting CLIMB mode, if within engagement parameters.

2-9-D-3.

APR 2 - NO DOPPLER

If doppler signal is lost, continue with APR 2 profile. If doppler is acquired during approach, doppler will synchronize and approach should continue normally. If approach is completed without doppler, helicopter will obtain airspeed between 0 and 40 KIAS when VEL HLD is captured, and lateral velocity up to 20 knots plus wind drift. Attempt to acquire doppler by beeping cyclic aft until airspeed goes to zero. If doppler is not acquired by this point, or visual contact is not made, initiate a go around. 7

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-69.4 2-9-D-4.

FAA APPROVED

MARK ON TARGET (MOT) NOTE

MOTmode may be engagedanywhere within APR 1 limits if doppler is valid. To reduce size of MOT pattern, MOT mode should be engagedat 200to 500 feet AGL and 60 knots groundspeed.

DPLR switch - NORM.

During final approachdescent:

AL-300- Set desired hover height between50 and 150feet AGL. EFISmode selector - HOV. Dopplerground speed - Monitor.

MOT light approach.

EADI annunciations - Check VEL HLD and RADALT ON.

MOTswitch - Press ON when directly over target. Verifylight illuminates.

Monitor descent and decelerationto 200feet AGL and 60 knots groundspeed. NOTE Hover waypoint and final approach course into wind will be established at MOT mode engagement. Wind will continue to be updated during approach until IAS drops below 30 knots on final approach. Below 30 KIAS,wind speed and direction display may be inaccurate. Note any changes in wind direction and speed prior to final approach. LNAVwind velocitywill be displayeddown to 5 knots,and EHSI wind vector and velocity will be displayeddown to 1 knot. When MOTis selected, helicopter will turn shortest direction downwind, then capture60 KIAS plus wind seenat start of approach. A patternwill beflown so that helicopter arrives at final descent point with 60 knot groundspeedat least 5 seconds prior to start of descent.

EHSI- Adjust rangeas desired to keep hover waypoint and final approach course In view. Mission equipment- Check.

8

Extinguished at completion of

NOTE

MOTmode will be cancelledby:

PressingMOTbutton on SARengage panel SelectingSBYthrough either MS-700 Flight Director ModeSelectoror Remote NAVSTBYbutton on either cyclic SelectingVEL HLDmode if below 60 knots airspeed SelectingRADALTmode SelectingAPR1,APR2,or CLIMBmode;if within specific engagementparameters for that mode.

2-9-D-5.

MARK ON TARGET (MOT)-NO DOPPLER

Valid doppler Is required to engage MOT mode. If doppler is lost, approach will continue unless doppler off time exceeds approximately 8 seconds in a turn, or 32 seconds level. MOT final hover position accuracywill be degradedwhen doppler is in memory. If doppler off time is exceeded, autopilot will disengage MOT mode and engage IAS and RADALT after rolling out on final approach. To complete approach, engage APR2 and refer to APR2 - NO DOPPLER.


2-9-D-6.

BHT-412-FMS-69.4

HOVER

IF HARDOVER OCCURS DURING AUTOMATIC APPROACH AND HOVER OPERATIONS, IMMEDIATELY TAKE OVER MANUALLY AND INITIATE CLIMB MODE. RETURN TO A SAFE ALTITUDE AND CORRECT MALFUNCTION PRIOR TO RESUMING AUTOMATIC APPROACH AND HOVER OPERATIONS. CAUTION

WHEN IN VEL HLD MODE, SELECTION OF ANY FLIGHT DIRECTOR MODES (PILOT OR COPILOT) WILL DISENGAGE VEL HLD.

DOPPLER IS REQUIRED FOR EXTENDED AUTOMATIC HOVER OPERATIONS. IF DOPPLER GOES INTO MEMORY, AUTOPILOT WILL MAINTAIN A CONSTANT VELOCITY SEEN AT TIME OF DOPPLER INMEMORY. IF PILOT HAS NO OUTSIDE REFERENCE WHEN IN ACCELEROMETER VELOCITY HOLD, THE CONDITION CANNOT BE MAINTAINED FOR EXTENDED PERIODS OF TIME WAITING FOR DOPPLER SIGNAL TO BE REACQUIRED. THE ALLOWABLE TIME PRIOR TO INITIATION OF CLIMB MODE WILL VARY DEPENDING ON CONDITIONS, BUT A MAXIMUM OF 45 SECONDS IN A STABLE HOVER AND 30 SECONDS IF MOVING AT ANY VELOCITY IS RECOMMENDED. IF PILOT IS UNSURE OF HELICOPTER'S ATTITUDE, AIRSPEED, OR DIRECTION; CLIMB MODE CAN BE ENGAGEDANYWHERE IN AUTOMATIC HOVER ENVELOPE BY PRESSING COLLECTIVE GO AROUND SWITCH.

MONITOR POWER DURING HOVER OPERATIONS. COLLECTIVE LIMIT IS 100% MAST TORQUE OR 810°C ITT. MANUALLY ADJUST COLLECTIVE WITH COLLECTIVE TRIM RELEASED, AS REQUIRED, TO MAINTAIN POWER WITHIN TRANSITORY LIMITS. DO NOT PUSH AGAINST COLLECTIVE TRIM OR A SLOW COLLECTIVE OSCILLATION MAY DEVELOP.

Coupled hover is normally engaged automatically at completion of APR 2 or MOT profile, but can be engaged manually by selecting RAD ALT and VEL HLD modes on autopilot controller. Hover velocities can be commanded by using cyclic beeper in desired direction up to 10 knots laterally and aft, and 60 knots forward. The velocity reference will be displayed on AL-300 for five seconds for direction actuated by cyclic beeper switch. This will be indicated by a minus sign for aft or left, and no sign for right or forward. Cyclic beeper switch - Select direction and velocity desired. Press center of beeper switch to return hover reference to zero. NOTE

If in IMC, a maximum of 5 knots laterally and aft is recommended, depending on gross weight and wind conditions. If doppler return conditions over water are marginal, maintaining hover velocities less than 5 knots will reduce the possibility of doppler memory excursions. Small hover corrections should be made by flying against trim, with zero velocity as the reference. Hover movements can be made by momentarily pressing cyclic FORCE TRIM release switch, but this is not recommended unless outside references are available. If cyclic FORCETRIM release switch Is pressed, a new hover velocity reference will be established at the time of FORCE TRIM switch release.

9


BHT-412-FMS-69.4

CLTV - YAW TRIM UP and DN switch desired. AL-300 desired.

Select

RAD ALT altitude,

As as

NOTE

Search and rescue operations will require extended hover in the Heightrange. AVOID Velocity Diagram Automatic hover operations between 50 and 150 feet AGL require that pilot be attentive to controls in the event of a sudden engine failure. Altitudes below 50 feet AGL can be selected using CLTV-YAW TRIM UP and DN switch. Below 100 feet AGL, pilot is required to have both hands lightly on controls to respond to engine failure or autopilot malfunction.

2-9-D-7.

CLIMB mode is selected by pressing collective GO AROUND switch, or CLIMB switch located on instrument panel. CLIMB mode may be selected anywhere in SAR mode envelope if RAD ALT is between 20 and 200 feet AGL and less than 60 KIAS. Initiation of CLIMB mode will return helicopter to 200 feet RAD ALT and 60 KIAS hold. NOTE

CLIMB mode may be initiated at anytime during SAR mode hover when doppler returns are lost. CLIMB mode will function with or without doppler inputs to autopilot.

Collective GO AROUND switch - Press.

Pilot HOIST EN switch - Press to transfer limited hover control to hoist operator (if desired). NOTE

Hoist operator has limited hover control by operating hover trim switch on hoist pendant in desired direction. White dot on HOIST EN switch will illuminate to indicate hoist operator is making control inputs. Helicopter can be flown up to 5 knots around a reference with hover velocity Pendant established by the pilot. switch is held in a given direction until desired speed is reached using outside references, then released to capture that velocity. To return to zero velocity, switch is operated in opposite direction until velocity is zero, then released. Pilot must monitor doppler bars to is responding helicopter ensure properly. Pilot may take hover trim control away from hoist operator by pressing HOIST EN switch, or simply by overriding hoist operator inputs, using cyclic stick or cyclic beep switch.

10

CLIMB

CAUTION

MONITOR POWER DURING INITIAL FORWARD TRANSITION. MANUALLY WITH COLLECTIVE ADJUST COLLECTIVE TRIM RELEASED AS REQUIRED TO MAINTAIN POWER WITHIN LIMITS. DO NOT PUSH AGAINST COLLECTIVE TRIM OR A SLOW COLLECTIVE OSCILLATION MAY DEVELOP. NOTE

When CLIMB mode is selected, helicopter will begin a slow, smooth transition into forward flight. Collective will maintain reference hover height acceleration, then during forward capture a 200 fpm climb rate.

EADI -

Check CLIMB annunciation. NOTE

When helicopter reaches 60 KIAS, IAS hold mode will engage. Upon reaching


FAA APPROVED

200 feet AGL, RAD ALT mode will modes are engage. When both engaged, CLIMB mode will extinguish.

During climb, VEL HLD may be engaged at anytime by momentarily pressing cyclic FORCE TRIM release switch or longitudinal beep switch. RAD ALT may be engaged at anytime by momentarily pressing CLTV-YAW TRIM REL switch or beeping CLTVYAW TRIM switch UP or DN.

During climb, roll axis will maintain zero lateral doppler velocity. Heading can be changed using CLTV-YAW TRIM switch in L or R position, and will

produce a flat uncoordinated turn to selected heading.

EADI mode.

Monitor for completion

of CLIMB

2-9-D-8. HOIST OPERATIONS

Refer to appropriate FMS and/or STC for approved limitations and procedures for hoist operations (if installed). HOIST PWR switch - ON.

Pilot HOIST EN switch - Press to transfer limited hover control to hoist operator (if desired).

11/12


3-1.

BHT-412-FMS-69.4

INTRODUCTION

For emergency procedures relating to optional installed equipment, refer to appropriate FMS or STC.

For emergency procedures relating to EFIS or AP, refer to BHT-412-FMS-38.4.

determine fault condition. Refer to Table 3-3 for list of error codes.

During SAR coupled mode operation, failure or abnormal disengagement of either AP1 or AP2 will not decouple the system. AFCS operation will revert to the remaining autopilot without interruption. Reengagement of the autopilot may be attempted without affecting mode operation.

THE FZ-706 FLIGHT CONTROL COMPUTER DOES NOT RECOGNIZE ENGINE FAILURE. FOLLOWING ENGINE FAILURE, COMPUTER WILL CONTINUE TO LIMIT REMAINING ENGINE TO APPROPRIATE LIMITS STATED IN SECTION 1 - LIMITATIONS, UNTIL PILOT DISENGAGES COLLECTIVE COUPLING OR MANUALLY OVERRIDES COLLECTIVE LIMITING. UNDER CERTAIN LOW OAT CONDITIONS, AFCS COMPUTER MAY EXCEED SINGLE ENGINE TORQUE LIMIT.

In the event of any succeeding failure of required equipment, the coupled SAR mode operation shall be terminated by the pilot. NOTE

The AFCS, if coupled in APR 1 mode, will remain coupled to heading mode following failure or abnormal disengagement of either AP1 or AP2.

3-15-A. SINGLE AFCS FAILURE *

INDICATIONS:

Appropriate AP FAIL light illuminated

3-15. SAR MODE MALFUNCTIONS

Automatic recouple of system to remaining AFCS.

*

All required equipment shall be operable prior to commencement of SAR modes APR 1, APR 2, MOT, VEL HLD, RAD ALT, or CLIMB. AP1 and AP2 shall be engaged in ATT mode. NOTE

If fault exists, use error code displayed on AL-300 Air Data Display to

PROCEDURE:

Attempt reengagement of affected system.

3-15-B. *

DUAL AFCS FAILURE

INDICATION:

AP1 and AP2 FAIL lights illuminated.

13

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-69.4 *

FAA APPROVED

will appear over RADALT display, indicating collective axis is no longer coupled to AFCS.

PROCEDURE:

NOTE

During SAR coupled mode operation, failure or abnormal disengagementof both AP1 and AP2 will decouple the system.

Loss of both AP1 and AP2 during APR 1 mode: SelectCLIMBmode. Manuallyfly helicopterto follow pitch and roll commandbars and collective cue. NOTE

At airspeedsabove 60 KIAS,command bars will provide cues for holding 70 KIASand 700fpm climb.

Loss of both AP1 and AP2 during APR 2 or VELHLD/RADALT modes: SelectCLIMBmode by pressing collective GOAROUNDswitch or instrument panelCLIMBswitch. Manuallyfly helicopterto follow pitch and roll command bars and collective cue. NOTE

At airspeedsbelow 60 KIAS,command bars will provide cues for holding 60 KIAS and 200 fpm climb to 200 feet AGL. Reengagementof AP1 or AP2 may be attempted after helicopter has been returned to safe flight conditions. Failedautopilot - Engage.

3-15-C. RAD ALT FAILURE

If coupled to failed RAD ALT, collective command cue will bias out of view. A red X

The pilot must assumecollective and altitude control by cross referencing the analog display.

Continuation of an uncoupled IMC approach is not recommended. 3-15-C-1.

SINGLE RAD ALT FAILURE

If cross-side RAD ALT is valid, collective may

be re-coupledby pressing FDSELswitch. 3-15-C-2.

DUAL RAD ALT FAILURE

and Manually fly collective/altitude discontinue SAR operations under IMC conditions.

3-15-D. WEATHER RADAR FAILURE

Referto WeatherRadar Operator'sManual for weather radar emergency and malfunction procedures(if installed).

3-15-E.

FLIGHT MANAGEMENT SYSTEM FAILURE

Flight Manual Refer to appropriate Supplement for installed Flight Management System(FMS).

3-15-F.

DOPPLER FAILURE

Doppler is required for extended automatic hover operations. If valid Doppler signal is lost, autopilot will maintain a constant velocity seen at time of signal loss. If pilot has no outside reference when in Accelerometer Velocity Hold, the condition cannot be maintainedfor extendedperiods of time waiting for Doppler signal to be reacquired. The allowable time, prior to initiation of Climb mode, will vary depending on conditions, but a maximumof 45 seconds

14


BHT-412-FMS-69.4

in a stable hover and 30 seconds if moving at any velocity is recommended. If pilot is unsure of helicopter's attitude, airspeed, or direction; Climb mode can be engaged anywhere in automatic hover envelope by pressing Collective Go Around switch.

Loss of Doppler, or valid Doppler information, will be displayed on EFIS Hover Mode by DOPP enunciator turning amber and/or loss of velocity bars.

Loss of Doppler valid signal will be displayed by FMS message. Table 3-3

PWR UP X X X X X X X X X X X X X X X X X

PRE FLT X X

CONT TEST X X

X X X X

X X X X

X

X X X X X

X X X X X X X X X X X X X X X X X

AL-300 DISPLAY EO El E2 E3 E4 E5 E6 E7 E8 E9 E10 Ell E12 E13 E14 E15 E16 E17 E18 E19 E20 E21 E22 E23 E24 E25 E26 E27 E28 E29 E30 E31 E32 E33 E34 E35 E36 E37 E38 E39 E40 E41


ERROR CODE DESCRIPTION LOST PROCESSOR RAM PARTITION FAIL INSTRUCTION SET FAIL REGISTERTEST FAIL RAM RD/WR FAIL EPROM CHECKSUM FAIL CPU FAIL D/A-A/D FAIL INTERRUPT REG FAIL +15V MONITOR FAIL (POSSIBLE HARDWARE E/A FAIL) -15V MONITOR FAIL +5V MONITOR FAIL +10V MONITOR FAIL LVC MONITOR FAIL SAWTOOTH MONITOR FAIL ACTUATOR REFERENCE MONITOR FAIL WATCHDOG MONITOR FAIL TICKET FAILURE LVC FAIL SAWTOOTH FAIL ACTUATOR REFERENCE FAIL WATCHDOG FAIL TRIM END AROUND FAIL TRMMON CKSUM FAIL NO YAW TRIM NO ROLL TRIM NO PITCH TRIM YAW TRIM FAIL ROLL TRIM FAIL PITCHTRIM FAIL YAW TRIM RUNAWAY ROLL TRIM RUNAWAY PITCHTRIM RUNAWAY TRMRUN CKSUM FAIL COLLECTIVE SERVO FAIL YAW SERIES SERVO FAIL ROLL SERIES SERVO FAIL PITCH SERIES SERVO FAIL

15


BHT-412-FMS-69.4

Table 3-3

PWR UP

ERROR CODE DESCRIPTION VG MISCOMPARE

X X X X X X X X X X X X X X X X X X X X X X X

AL-300 DISPLAY E42 E43 E44 E45 E46 E47 E48 E49 E50 E51 E52 E53 E55 E56 E101 E102 E103 E104 E105 E106 E107 E108 E109 E110 E111 E112 E113 E114 E115 E116 E117 E118 E119 E120 E121 E122 E123

X X X X X X

E125 E126 E127 E128 E129 E130

POD DURING TEST ROD DURING TEST YOD+YFTR DURING TEST COD DURING TEST FTR DURING TEST CFTR DURING TEST

PRE FLT

X

X X X X X X X

X X X 16


CONT TEST X X X X X X X

E124

ON-SIDE/CROSS-SIDE TORQUE INPUT OPEN ON-SIDE/CROSS-SIDE TEMP INPUT OPEN OVERTORQUE AND COLL ON OVERTEMP AND COLL ON

RADALT COMPARISON FAIL VRTACC COMPARISON FAIL COLLECTIVE CMT FAIL ERROR CODE STORAGE ALERT: <100 AVAILABLE ERROR CODE STORAGE ALERT: NONE AVAILABLE PITCH ARL TEST FAIL ROLL ARL TEST FAIL PITCH ACT MOVED DURING ARL TEST ROLL ACT MOVED DURING ARL TEST PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED YAW TRIM ACT MOVED NO YAW TRIM NO ROLL TRIM NO PITCH TRIM YAW TRIM BACKWARDS ROLL TRIM BACKWARDS PITCH TRIM BACKWARDS SERVO VALID DELAY (2 SEC) FAIL WATCHDOG CANNOT DISENGAGE AP YAW ACT/SERVO AMP FAIL ROLL ACT/SERVO AMP FAIL PITCH ACT/SERVO AMP FAIL COLL ACT/SERVO AMP FAIL FD ANN FAIL COLL RELAY INOPERATIVE COLL INHIBIT INOPERATIVE

FDANN FAILOTHER

E131

TRIM SWITCHOFFDURINGTEST

E138

YAWTRIMACTMOVED

E132 E133 E134 E135 E136 E137 E139 E140

COLL TRIM SWITCH OFF DURING TEST PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED YAW TRIM ACT MOVED PITCH TRIM ACT MOVED ROLL TRIM ACT MOVED YAW SERIES ACT MOVED COLL ACT MOVED


BHT-412-FMS-69.4

Table 3-3

PWR UP

PRE FLT X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X

CONT AL-300 TEST DISPLAY E141 E142 E143 E144 E145 E146 E147 E148 E149 E150 E151 E152 E153 E154 E155 E156 E157 E201 E202 E203 E204 E205 E206 E207 E208 E209 E210 E211 E212 E213 E214 E215 E216 E217 E218 E219 E220 E221 E222 E223 E224 E225 E226 E227 E228 E229 E230 E231 E232 E233 E234


ERROR CODE DESCRIPTION PITCH TRIM (H) E/A ON PITCH TRIM (L) E/A ON ROLL TRIM (H) E/A ON ROLL TRIM (L) E/A ON YAW TRIM (H) E/A ON YAW TRIM (L) E/A ON TRIM ENABLE FAIL TRIM ONE-SHOT FAIL PITCH TRIM (H) E/A OFF PITCH TRIM (L) E/A OFF ROLL TRIM (H) E/A OFF ROLL TRIM (L) E/A OFF YAW TRIM (H) E/A OFF YAW TRIM (L) E/A OFF SERVO VALID FAIL TRIM END AROUND FAIL APON OTHER INPUT FAIL AL-300 KNOB MOTION AIRSPEED MORE THAN 60 KNOTS PRIMARY RADALT MORE THAN 20 FT ANALOG GROUND FAIL LATERAL ACCELEROMETER FAIL LONGITUDINAL ACCELEROMETER FAIL VERTICAL ACCELEROMETER FAIL ROLL ATTITUDE MORE THAN 10 DEG PITCH ATTITUDE MORE THAN 10 DEG YAW RATE GYRO MORE THAN 10 DEG/SEC DIRECTIONAL GYRO FAIL VERTICAL GYRO NOT VALID DIRECTIONAL GYRO NOT VALID PRIMARY RADIO ALTIMETER NOT VALID AIR DATA SENSOR NOT VALID SERIAL IO NOT VALID OTHER AUTOPILOT NOT IN TEST MODE SELECTOR BUTTON(S) PRESSED AP CONTROLLER BUTTON(S) PRESSED SAR CONTROLLER BUTTON(S) PRESSED HARDWARE NOT VALID FOUR-WAY BEEP SWITCH PRESSED ENC ALT LESS THAN -600 FT OR GREATER THAN 12,000 FT BARO ALT LESS THAN -600 FT OR GREATER THAN 12,000 FT POD DURING TEST ROD DURING TEST YOD+YFTR DURING TEST COD DURING TEST FTR DURING TEST CFTR DURING TEST TRIM SWITCH OFF DURING TEST COLL TRIM SWITCH OFF DURING TEST ENG TORQUE LESS THAN 5% OR GREATER THAN 30% ENG TEMP LESS THAN 150 DEG OR GREATER THAN 900 DEG C

17


BHT-412-FMS-69.4

Table 3-3

PWR UP

18

PRE FLT X X X X X X X X X X X X X X

CONT AL-300 TEST DISPLAY E235 E236 E237 E238 E239 E240 E241 E242 E243 E244 E245 E246 E247 E248



AIRSPEED INPUT OPEN HEADING NOT VALID ENCODING ALTIMETER NOT VALID DOPPLER ARINC 429 BUS NOT VALID LNAV ARINC 429 BUS NOT VALID C/A RESET BUTTON PRESSED CYCLIC BEEP ACTIVE ON DURING TEST ERASE EEPROM SWITCH ON DURING TEST READ EEPROM SWITCH ON DURING TEST VELHLD SYNC SWITCH ON DURING TEST SECONDARY RADALT MORE THAN 20 FEET SECONDARY RADIO ALTIMETER NOT VALID


FAA APPROVED

WARNING

DUE TO SAFETY MARGIN DESIGNED INTO COLLECTIVE COUPLING

SYSTEM, FLIGHT MANUAL PERFORMANCE MAY NOT BE OBTAINED DURING THREE CUE COUPLED MODE.

19



20

T N I Section 1 R P N O R I O CT F U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

1-1.

INSTRUMENT PANEL AND CONSOLES

BHT-412-FMS-69.4

1-3.

1-3-A.

ELECTRICAL SYSTEMS

DC ELECTRICAL SYSTEM


Refer to Tables 1-1 and 1-2.

Figures 1-1 and 1-2 provide instrument panel and pedestal.

1-3-B. AC ELECTRICAL SYSTEM Refer to Table 1-5.

21

BHT-412-FMS-69.4 MANUFACTURER'S DATA

T N I R P N O R I O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O


BHT-412-FMS-69.4

LOCATION:FMS may be in either location.(typical)

412FS69.4-1-2

Figure 1-2. Pedestal

23

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTUREER'SDATA

BHT-412-FMS-69.4

Table 1-1. DC electrical system bus loading (Nonessential buses)


(For complete list, see BHT-412-FM-4/FMS-38.4)

NONESSENTIAL DC BUS 2 DF (Receiver) DATA NAV DPLR

NOTES: 1. Automatically sheds with one generator inoperative, but can be reenergized by placing the NON-ESNTL BUS switch in the MANUAL position. 2.

Automatically sheds for 30 and 90 minute flight with electrical power supplied by the battery only.

Table 1-2. DC electrical system bus loading (Essential buses)

ESSENTIAL DC BUS 1

(For complete list, see BHT-412-FM-4/FMS-38.4)

ESSENTIAL DC BUS 2

NAV-COM LNAV/FMS/RRS/ADC

NOTE: 1. Sheds for 90 minute flight with electrical power supplied by battery only, and with emergency load switch in EMERG LOAD position.

Table 1-5. AC electrical system bus loading

NO. 1 INVERTER (For complete list, see BHT-412-FM-4/FMS-38.4)

24

26 VAC BUS 2 DF (Receiver)

T N I BELL 412 R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-70.4

MODEL


FLIGHT MANAGEMENT SYSTEM UNS-1 C/D 412-899-381

(S/N 36218 AND SUB) CERTIFIED 29 JANUARY 1999

This supplement shall be attached to Bell Helicopter Model 412EP Flight Manual when Universal Avionics Flight Management System is installed in accordance with Drawing 412-899-381.

Information contained herein supplements information of Basic Flight Manual. For Limitations, Procedures, and

Performance Data not contained in this supplement, consult Basic Flight Manual.

NOTICE COPYRIGHT COPYRIGHT

Bell

2000

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

POST OFFICE BOX 402

HelicopterTEXTR

A Subsidiaryof Textron Inc

FORT

WORTH, TEXAS 76101

29 JANUARY

REISSUE -

1999

27 APRIL 2000


NOTICE PAGE


Manufacturer's Data portion of this supplement is proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for

any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron Inc.


NP


LOG OF REVISIONS

.................. 29 JAN 99 Original ...................... Reissue ...................... 0 .................. 27 APR 00

LOG OF PAGES

REVISION NO.

PAGE

FLIGHT MANUAL

Title .......................................................... NP ............................................................ A- B ...................................................... 1/11............................................................. 1- 6 ........................................................

REVISION NO.

PAGE

MANUFACTURER'S DATA

0 0 0 0 0

7/8 ............................................................

0

NOTE



FAA APPROVED


Original . ......... 0 ................... 29 JAN 99 Reissue ...................... 0 .................. 27 APR 00

APPROVED

MANAGER


B

DATE

APR2 7 2000

BHT-412-FMS-70.4


Customized installation of universal Avionics

HF

High Frequency

comprises the following components:

ILS IFR

Instrument Landing System Instrument Flight Rules

INT LDA

Integrity Localizer Type Directional Aid

LNAV LOC/BC

Long Range Navigation Localizer Backcourse

MAP

Missed Approach Procedure

MLS

Microwave Landing System

MSG

Message

NAV

Navigation

NDB

Nondirectional Beacon

UNS-1C/D Flight Management system * Flat Panel Control Display Unit * Navigation Management Computer * GPS Antenna

ABBREVIATIONS

APR

Approach

CDU

Command Display Unit

DF

Direction Finder

DME

Distance Measuring Equipment

NM

Nautical Miles

EADI

EFIS Attitude Director Indicator

NMU RAIM

EFIS

Electronic Flight Instrument System

RRS

Navigation Management Unit Receiver Autonomous Integrity Monitoring Radio Reference Sensor

RNAV

Area Navigation

SAR

Search and Rescue

SX

Selected Cross Track

EHSI ELT

EFIS Horizontal Situation Indicator Emergency Locator Transmitter

FAF

Final Approach Fix

TACAN

Tactical Air Navigation

FD

Flight Director

UHF

Ultra High Frequency

FMS

Flight Management System

VHF

Very High Frequency

FPCDU

Flat Panel Control Display Unit

VOR

VHF Omnidirectional Range

GPS


WPT

Waypoint

HDG

Heading

i/ii


FAA APPROVED

Section

1-1.

1

INTRODUCTION

1-5.

CONFIGURATION

Limitations described in this supplement are in addition to those described in Basic Flight

UNS-1C/D is approved

Manual and apply only to Model 412EP

operations,when equippedwith the following

helicopter equipped with Universal Avionics

kits:

for IMC en route

UNS-1C/D Flight Management System

EFIS (412-705-009-101 /-105) using symbol generator (P/N 7011672-708). Refer to BHT-412-FMS-38.4.

installed. 1-1-A.

LONG RANGE NAV

Software number of Operators' Manual shall

match VERSION number annunciated on

or:

Initialization page.

Universal Avionics

SAR modes (412-705-008). Refer to BHT-412-FMS-39.3/39.4 or BHT-412-

(UNS-1C/D) Operator's

Manual (Publication Number 2423sv703),

dated 16 FEB 98, and Operators Checklist (Publication Number 2411sv703), dated 25 JUL 97, shall be immediately available to flight crew whenever navigation is predicated on use of system.

1-3.

FMS-69.4and BHT-412-FMS-38.4.

UNS-1C/D is approved for IMC GPS enroute and non-precision approach operations when equipped with the following kits: EFIS (412-705-009-103) using symbol generator (P/N 7011672-709).

TYPES OF OPERATION

UNS-1C/D is approved for IMC and GPS nonprecision en route operations, when equipped as outlined in Section 1-5.

IFR en route and terminal navigation is prohibited, unless pilot verifies currency of database or verifies each selected waypoint for accuracy by reference to current approved

or:

SAR modes (412-705-008). Refer to

BHT-412-FMS-39.3/39.4or BHT-412FMS-694 and BHT-412-FMS-38.4.

UNS-1C/D can be located in left or right center

data.

console for Search and Rescue (SAR)

Instrument and non-precison approaches

operation (kit 412-705-008).

must be conducted in approach mode and RAIM must be available at Final Approach Fix (FAF). Only VOR-DME, GPS, NDB, and VOR are permitted under IMC conditions. Displayed UNS-1C/D navigation parameters are referenced to Magnetic North. Operation of helicopter is limited to latitudes between 73 degrees North and 60 degrees South, unless

For Single Pilot GPS non-precision approach operations, UNS-1C/D shall be located in right center console. UNS-1C/D shall also have an operable Flight Director (FD). System shall utilize Software Number 703. Helicopter

shall

have other

approved

magnetic variation is manually entered by

navigationequipment,appropriate to route of

pilot.

flight, installed and operational. 1

BHT-412-FMS-70.4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 1-6.

WEIGHT AND CENTER OF

1-20.

GRAVITY

Refer to Figure 1-6 for placards.


2

PLACARDS


FAA APPROVED

Location: Varies per installation (typical)

NOTICE:

THIS PLACARD IS APPLICABLE ONLY

WHEN EFIS SYMBOL GENERATOR (P/N

7011672-708) IS INSTALLED.

Figure 1-6 Placards 3

T N Section2 I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O BHT-412-FMS-70.4

2-6.

FAA APPROVED

SYSTEMS CHECK

points and/or by using other navigation equipment such as VOR, DME, NDB, and/or radar.

FMS CHECK

2-6-A.

FMSswitch- ON.

2-9.

During Self-Test, verify FMS annunciators (HDG, SX, and RAIM) are illuminated and following indications are displayed on EHSI. Distance

123.4 NM

Ground speed

250 knots

Wind speed

50 knots

After power-on

self-tests

have been


Normal operating procedures are outlined in Universal UNS-1C/D Operators Manual, (Publication Number 2423sv703) and UNS-1C/ D Operator's Checklist (Publication Number 2411sv703).

CHECK VALIDITY OF

UTILIZED

completed, INITIALIZATION page will appear. Expiration date of database will appear at bottom of display. If actual date is later than expiration date, UNS-1C/D will notify pilot with a DATABASE EXPIRED message. Refer to

2-9-A.

operators manual for navigation database update Instructions.

Refer to Universal Avionics (UNS-1C/D)

Doppler (if installed) - ON, mode/s

MAGNETIC VARIATION IF UNS-1C/D IS TO BE USED WITH LATITUDES IN EXCESS OF 60° . SYSTEM ANNUNCIATORS

Operator's Manual.

selected.

Following two annunciators are interfaced

EHSI course display - Set to LNAV.

with EFIS.

UNS-1C/D system position shall be checked

Waypoint -

WPT (amber) annunciation

on

for accuracy prior to use as a means of

pilot and copilot EHSI.

conditions:

seconds prior to waypoint in en route mode, and five seconds prior to approach mode.

navigation, and under the following Following a period of POSITION UNCERTAIN dead reckoning.

When a crosscheck with other approved onboard navigation equipment reveals an error greater than 3 NM.

4

Meaning - Indicates waypoint alert 15

Message - FMS MSG (amber) annunciation EADI. on on pilot pilot and and copilot copilot EADI.

Meaning - Indicates a message is to be acknowledged on MSG page. For additional condition/modes

Helicopterposition should be verified

annunciations on instrument panel, see

by visually sighting ground reference

Section 3.

BHT-412-FMS-70.4

FAA APPROVED

T N I R P N O R I O T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O 2-9-B.

FLIGHT MANAGEMENT SYSTEM (FMS)

For IFR en route operations, verify approved,

current program/data base appears on page. Refer to UNS-1C/D initialization Operators Manual for currently installed Flight Management System database.

Terminal and approach database information

limited to VMC use only when EFISsymbol generator (P/N7011672-708) is installed. 2-9-B-1.

EFIS INTERFACE

Course deviation indicator on EHSI provides center to full scale offset of ± 5.0 nautical

miles crosstrack displacement for en route

upper right corner of EHSI. However, bearing pointer will not indicate waypoint bearing unless LNAV is selected on DC-811 Display Controller

with bearing pointer knob(s).

TO/FROM flag on EHSI is activated when a valid leg has been programmed. The legend changes from TO to FROM when passing the TO waypoint.

Master warning of system failure, dead reckoning, or no-valid leg is indicated by

displaying

a red X across the CDI on EHSI.

the MSG (message) light will Additionally, Additionally, the MSG (message) light will flash on the CDU/EADI. When the MSG key is activated, a sensor message page will appear,

mode.

indicating the occurrence.

When symbol generator (P/N 7011672-708) is installed, normal terminal/approach and FMS mode (GPS data base) terminal/approach course deviation is ± 1.2 nautical mile.

2-9-B-2.

When symbol generator (P/N 7011672-709) is

installed,

and FMS (GPS data base) is

selected, terminal mode course deviation is ± 1.0 nautical mile and final approach course deviation is ± 0.3 nautical mile.

AUTOPILOT COUPLING

LNAV/FMS may be coupled through autopilot when: FMS is selected as navigation source

on EFIS.

NAV is selected on FD mode selector. NOTE

NOTE

When LNAV is selected as navigation source, FMS (P/N 7011672-708)or LNAV (P/N 7011672-709) will be displayed in

When programming course segment ° changes greater than 90 , waypoint must be designated as OVFLY (overfly), if overstation passage is desired.

5

BHT-412-FMS-70.4

FAA APPROVED

T N Section3 I R P N O R I O T F C U T O OD N R E P E E R R C S G Section 4ON N I R W O E I F V NLY Section 5 O 3-15.

CAUTION AND

ADVISORY ANNUNCIATOR

Table 3-2.

Caution (Amber) / Advisory (Green) Lights

PANEL WORDING FMS RAIM (amber)

FAULT CONDITION

CORRECTIVE ACTION

GPS RAIM not available.

SX (green)

Pilot selected CROSSTRACK course. Pilot selected COMMAND HEADING

IFR en route: Continue to navigate using UNS-1C/D FMS, while verifying position using another IFR approved navigation system or revert to alternate means of navigation. Informational.

HDG (green)

Informational.


No change from basic manual. 6

T N I R P N O . OR I T F C U T O OD N R E P E E R R C S N G O N I R W O E I F V NLY O MANUFACTURER'S DATA

Section

BHT-412-FMS-70.4

1

1-18. NAVIGATION SYSTEMS AND

INSTRUMENTS

1-18-D.

UNS-1C/D (GPS) LONG RANGE NAV/FLIGHT MANAGEMENT SYSTEM

Global Positioning System (GPS) is global

coverage satellite based navigation system. It provides Standard Positioning Service (SPS) to all equipped users. Universal Avionics UNS-1C/D is a long range

GPS/DME based RNAV and Flight Management System (FMS) consisting of a Dzus-rail mounted receiver/ display unit and antenna. This system provides pilot with

navigational guidance and position

information in three dimensions: latitude,

longitude and altitude. Navigational information is contained in Internal database.

In addition to Waypoint (WPT) and FMS Message (MSG) annunciator located on EHSI display, Universal Avionics UNS-1C/D system

has three external annunciators located on instrument panelsas follows: HDG - Heading

SX - Selected Cross Track RAIM - Integrity.

Display of FMS navigational data and guidance is made on EFIS system by selecting LNAVon EFIScontroller(s).

7/8

BHT-412-FMS-72.4

T N I ROTORCRAFT R FLIGHT MANUALN P O R I O CT F COCKPIT VOICE RECORDER U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O SUPPLEMENT

412-705-001-115 412-704-026-101 CERTIFIED 23 MAY 2002

This supplement shall be attached to the Bell Helicopter Model 412 Flight Manual when the 412-705-001-115 Cockpit Voice Recorder kit or the 412-704-026-101 Cockpit Voice Recorder Retrofit kit is installed.

Information contained herein supplements and supersedes information in the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, refer to the basic Flight Manual.

COPYRIGHT NOTICE 2002 COPYRIGHT BELL ® HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON CANADA LTD. ALL RIGHTS RESERVED

POST OFFICE BOX 482


23 MAY 2002

BHT-412-FMS-72.4

NOTICE PAGE

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O PROPRIETARY RIGHTS NOTICE

Manufacturer’s Data portion of this supplement is proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron Inc.

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 76101-0482 NP

23 MAY 2002

BHT-412-FMS-72.4

LOG OF REVISIONS Original ................... 0 ......................23 MAY 02

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O LOG OF PAGES

PAGE

REVISION NO.

Title............................................................ 0 NP .............................................................. 0 A/B............................................................. 0 C/D............................................................. 0 i/ii ............................................................... 0 1 — 2 ......................................................... 0 3/4 .............................................................. 0

PAGE

REVISION NO.

MANUFACTURER’S DATA 5 — 6 ......................................................... 0

NOTE Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages. 23 MAY 2002

A/B

BHT-412-FMS-72.4

LOG OF FAA APPROVED REVISIONS Original ................... 0 ......................23 MAY 02

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O APPROVED:

DATE:

MANAGER

ROTORCRAFT CERTIFICATION OFFICE FEDERAL AVIATION ADMINISTRATION FT. WORTH, TX 76193-0170 23 MAY 2002

C/D

BHT-412-FMS-72.4

GENERAL INFORMATION The Cockpit Voice Recorder kit consists of a Cockpit Voice Recorder (CVR) unit, installed in the nose lower shelf, and a Recorder Control Unit (RCU) located in the center pedestal. This kit is designed to be utilized with aircraft headsets that use dynamic microphones, which provide continuous live microphone audio to the cockpit voice recorder. Because the 212-077-200-( ) audio system does not provide microphone bias voltage when not keyed, the use of other types of microphones requires FAA approval.


The Cockpit Voice Recorder has the capability to record two hours of voice data, stored on a crash-survivable solid-state memory. Four channels are provided, three with narrow audio bandwidth, and one with a high bandwidth for use by the Cockpit Area Microphone (CAM). Two of the narrow bandwidth channels are used to record the pilot’s and copilot’s hot mic/headset audio. The other narrow bandwidth channel is used to record the time correlation signal from the FDR (if installed). Crash Protected Memory (CPM) is designed to retain, without loss or alteration, stored information without applied power for two years. The unit has an external Underwater Locator Beacon (ULB). The ULB is a self-contained locating device that aids in the underwater recovery of the cockpit voice recorder system. It automatically activates upon immersion in water (fresh and salt) and sends out periodic acoustic pulses for a minimum of 30 days. The front face of the RCU contains a Cockpit Area Microphone (CAM), a CVR TEST pushbutton and an amber CVR FAIL annunciator. It also includes a CVR ERASE button, which is deactivated for this installation, and a maintenance headset jack for monitoring the data being recorded by the Cockpit Voice Recorder. An inertia switch is installed in series with the Cockpit Voice Recorder power input to prevent overwriting of the recording after an accident. It is a full three-axis switch, mounted on the right hand side of the lower avionics shelf in the aircraft nose.

23 MAY 2002

i/ii

FAA APPROVED

BHT-412-FMS-72.4

Section 1 LIMITATIONS

T N I R P N O R I O CT F U T D O O N N R E P Section 2 E E R R C S N G O N I R W O E I F V NLY O 1

1-5-B.

OPTIONAL EQUIPMENT

COCKPIT VOICE RECORDER

Only headsets with a dynamic microphone are approved for use with the Cockpit Voice Recorder. Cockpit Voice Recorder operation with other types of microphones requires FAA approval.

NORMAL PROCEDURES

2

2-7.

BEFORE TAKEOFF

COCKPIT VOICE RECORDER CHECK Verify CVR FAIL annunciator — Not illuminated or blinking.

Release CVR TEST pushbutton. Verify CVR FAIL annunciator — Remains illuminated for more than 1 second and extinguishes within 5 seconds. Verify area microphone is unobstructed.

Press and hold CVR TEST pushbutton on RCU. Verify CVR FAIL annunciator — Illuminated.

23 MAY 2002

1

BHT-412-FMS-72.4

FAA APPROVED

Section 3 EMERGENCY/MALFUNCTION PROCEDURES

3


3-16. COCKPIT VOICE RECORDER

3-16A. COCKPIT VOICE RECORDER FAILURE — AT START-UP • INDICATION:

Cockpit Voice Recorder fails prestart check.

• PROCEDURE:

Refer to local operational or regulatory requirements.

2

23 MAY 2002

3-16B. COCKPIT VOICE RECORDER FAILURE — IN-FLIGHT • INDICATION:

Amber CVR FAIL annunciator illuminates during flight.

• PROCEDURE:

Maintenance action required.

FAA APPROVED

BHT-412-FMS-72.4

Section 4 PERFORMANCE

T N I R P N O R I Section 5 O T F C U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O 4


WEIGHT AND BALANCE

5


23 MAY 2002

3/4

MANUFACTURER’S DATA

BHT-412-FMS-72.4

Section 1 SYSTEM DESCRIPTION

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O 5

1-16. EMERGENCY EQUIPMENT COCKPIT VOICE RECORDER

The cockpit voice recorder Recorder Control Unit (RCU) consists of the following: • CVR TEST pushbutton — When depressed initiates the CVR self-test.

• CVR ERASE pushbutton — Deactivated in this installation.

Table 1-1.

• CVR FAIL annunciator (amber) — Provides indication of cockpit voice recorder status. • PHONES — Headphone monitor jack allows monitoring of mixed audio supplied to cockpit voice recorder using a 600-ohm headset. • MIC — Area microphone used to capture cockpit audio. Must not be covered by loose objects during all flight phases.

Cockpit Voice Recorder Circuit Breaker EMERG BUS

2337CB1 CVR

Disconnect No. 1 and No. 2 Main DC Buses to the Cockpit Voice Recorder

Location – Nose Electrical Compartment – LHS

23 MAY 2002

5

BHT-412-FMS-72.4

MANUFACTURER’S DATA

T N I R P N O R I O CT F U T D O O N N R E P E E R R C S N G O N I R W O E I F V NLY O Figure 1-4.

6

23 MAY 2002

Cockpit Voice Recorder – RCU

Bht-412-Fm-4 Bell 412ep Flight Manual

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