 407 - Flight Manual
• 407 Flight Manual 1 - Temporary Revisions
 Log of Temporary Revisions (16 Nov 2006)
 TR-9 - Sustained Hover and Vertical Takeoff/Landing Operations With
Tailwind (15 Jan 2002)
 TR-10 - Incorporation of Oil Cooler Blower Inlet Ducts and Bearing
Airflow Shields (Rev 1 - 25 Jul 2002)
 TR-11- FADEC Software Version 5.356 (16 Nov 2006)
 407-FM-1 - Flight Manual (Revision 7 - 30 Jul 2008)
 Front Matter
 Section 1 - Limitations
 Section 2 - Normal Procedures
 Section 3 - Emergency/Malfunction Procedures
 Section 4 - Performance
 Section 5 - Weight and Balance
 Appendix A - Optional Equipment Supplements
 407-FM-1-ATO - Title Page for Republic of Philippines Registered Helicopters
(Basic Issue: 12 Nov 1996)
 407-FM-CTA - Title Page for Brazil Registered Helicopters (Basic Issue: 03 Apr
1998)
 407-FM-IAC AR - Title Page for Interstate Aviation Committee - Aviation
Register Commonwealth of Independent States (Basic Issue: 20 May 1999)
 407 - Flight Manual Supplement
 407-FMS-1 - Lightweight Emergency Flotation Landing Gear (Basic Issue: 11
Apr 1996)
 407-FMS-2 - High Skid Gear (Basic Issue: 14 Feb 1996)
 407-FMS-3 - Particle Separator (Reissue - 16 Dec 2002)
 407-FMS-4 - Snow Deflector (Reissue - 16 Dec 2002)
 407-FMS-5 - Cargo Hook (Revision 1 - 25 Mar 2008)
 407-FMS-6 - Auxiliary Fuel Kit (Basic Issue: 20 Mar 1996)
 407-FMS-7 - Litter Kit (Revision 1 - 16 Sep 1999)
 407-FMS-17 - Cargo Tiedown Provisions Kit (Basic Issue: 01 Apr 1996)
 407-FMS-20 - KLN 89B GPS Navigator (Revision 1 - 26 Nov 1996)
 407-FMS-21 - Fire Detection System (Reissue - 07 Jul 2004)
 407-FMS-22 - Auxiliary Vertical Fin Strobe Lights (Basic Issue: 10 May 1996)
 407-FMS-23 - RYAN Traffic Collision Avoidance Device (Basic Issue: 15 May
1996)
 407-FMS-25 - Quiet Cruise Mode (Revision 1 - 30 Jul 2008)
 407-FMS-28 - Increased Internal Gross Weight (Reissue - 16 Dec 2002)
 407-FMS-31 - Increased APC Starter Generator Load Kit (Basic Issue: 15 Jun
2005)
 407-FMS-CAA - United Kingdom Registered Helicopters (Basic Issue: 08 Jan
2002)
 407-FMS-IAC AR - Interstate Aviation Committee - Aviation Register
Commonwealth of Independent States (Basic Issue: 20 May 1999)
 407 - Manufacturer's Data
 407-MD-1 - Manufacturer's Data (Revision 4 - 30 Apr 2008)
 Front Matter
 Section 1 - Systems Description
 Section 2 - Handling and Servicing
 Section 3 - Conversion Charts and Tables
 Section 4 - Expanded Performance
BHT-407-FM-1
LOG OF TEMPORARY REVISIONS
TEMP. REV. NO.
TITLE
DATE ISSUED
DATE
CANCELED
BHT-407-FM-1
FADEC Fault Annunciation
Interpretation
Revision 1
03 December 1996
15 June 2000
BHT-407-FM-1
NP Overspeed Trip Increase
03 December 1996
15 June 2000
BHT-407-FM-1
Airspeed Change to 125 KIAS and
Temporary Pedal Stop
16 December 1998
10 March 1999
BHT-407-FM-1
FADEC Software Version 5.202
22 December 1998
17 December 2002
BHT-407-FM-1
Hover Performance Correction for
Temporary Tail Rotor Pedal Stop
10 March 1999
17 December 2002
BHT-407-FM-1
VNE Increase to 130 KIAS
Reissue
3 June 1999
17 December 2002
BHT-407-FM-1
FADEC Direct Reversion to Manual
System, ASB 407-99-31
4 June 1999
17 December 2002
BHT-407-FM-1
VNE Increase to 140 KIAS
Revision 1
27 June 2000
17 December 2002
BHT-407-FM-1 (TR-9)
Sustained Hover and Vertical Takeoff/
Landing Operations with Tailwind
15 January 2002
BHT-407-FM-1 (TR-10)
Incorporation of Oil Cooler Blower Inlet
Ducts and Bearing Airflow Shields
Revision 1
25 July 2002
BHT-407-FM-1 (TR-11)
FADEC Software Version 5.356
(Reversionary Governor)
16 November 2006
This Log of Temporary Revisions provides the current status of each Temporary Revision issued against
the basic Flight Manual. It should be inserted at the back of the Flight Manual binder for quick and easy
reference.
16 November 2006
BHT-407-FM-1
ROTORCRAFT
FLIGHT MANUAL
TEMPORARY REVISION
FOR
SUSTAINED HOVER AND VERTICAL
TAKEOFF/LANDING OPERATIONS
WITH TAILWIND
Insert these Temporary Revision pages next to like-number
pages in the basic Flight Manual.
DO NOT remove existing pages. DO NOT remove temporary
pages until replacement pages are received or temporary
pages are canceled.
NOTE
For tracking purposes, Temporary Revisions are now being
numbered. This Temporary Revision is issued as TR-9,
following previously issued Temporary Revisions.
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
COPYRIGHT NOTICE
2002
COPYRIGHT
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESER
RESERVED
Bell Helicopter
A Subsidiary of Textron Inc.
POST OFFICE BOX 482
FORT WORTH, TEXAS 76101
09 FEBRUARY 1996
TEMPORARY REVISION (TR-9) — 15 JANUARY 2002
BHT-407-FM-1
LOG OF REVISIONS
Temporary (TR-9)...............................15 JAN 02
LOG OF PAGES
PAGE
REVISION
NO.
PAGE
REVISION
NO.
Title ........................................ Temporary (TR-9)
A ............................................. Temporary (TR-9)
C-D ......................................... Temporary (TR-9)
1-5 .......................................... Temporary (TR-9)
1-7 .......................................... Temporary (TR-9)
NOTICE
CHANGED INFORMATION IN THIS TEMPORARY REVISION APPEARS IN BOLD TYPE.
ALL OTHER INFORMATION IS OUTLINED.
TEMPORARY REVISION (TR-9) — 15 JAN 2002
A
BHT-407-FM-1
LOG OF TC APPROVED REVISIONS
Temporary (TR-9)...............................15 JAN 02
APPROVED
DATE
CHIEF, FLIGHT
FLIGHT TEST
TEST
CHIEF,
FOR
FOR
DIRECTOR —
— AIRCRAFT
AIRCRAFT CERTIFICATION
CERTIFICATION BRANCH
DIRECTOR
DEPARTMENT
OF
TRANSPORT
TRANSPORT CANADA
TEMPORARY REVISION (TR-9) — 15 JAN 2002
C
BHT-407-FM-1
LOG OF FAA APPROVED REVISIONS
Temporary (TR-9) .............................. 18 JAN 02
D
TEMPORARY REVISION (TR-9) — 15 JAN 2002
TC APPROVED
BHT-407-FM-1
Maximum allowable airspeed for sideward
and rearward flight or crosswind hover is
35 KTAS.
Sustained hover and vertical takeoff/
landing operation (greater than one minute)
with tailwind (relative winds within ± 90° of
tail) greater than 5 knots is prohibited.
1-8.
ALTITUDE
Maximum operating altitude is 20,000 feet Hp .
TEMPORARY REVISION (TR-9) — 15 JAN 2002
1-5
TC APPROVED
1-13-G.
BHT-407-FM-1
ENGINE OIL TEMPERATURE
Continuous operation
0 to 107°C
Maximum
107°C
CAUTION
IF HOVERING WITH A TAILWIND
GREATER THAN 5 KNOTS AT
OAT ABOVE 24°C (75°F),
CLOSELY MONITOR ENGINE
AND TRANSMISSION OIL
TEMPERATURES. IF ENGINE
OR TRANSMISSION OIL
TEMPERATURES
RISE
ABNORMALLY, TURN INTO WIND,
REDUCE POWER OR TRANSITION
TO FORWARD FLIGHT UNTIL
TEMPERATURE DECREASES.
NOTE
Positive temperature indication is when
the second segment of the trend arc is
illuminated.
1-14.
TRANSMISSION
TEMPORARY REVISION (TR-9) — 15 JAN 2002
1-7
BHT-407-FM-1
ROTORCRAFT
FLIGHT MANUAL
TEMPORARY REVISION FOR THE
INCORPORATION OF OIL COOLER BLOWER
INLET DUCTS AND BEARING
AIRFLOW SHIELDS
This Temporary Revision supersedes and replaces in its entirety, Temporary
Revision for Sustained Hover and Vertical Takeoff/Landing Operations with
Tailwind, TR-9 dated 15 January 2002, when Oil Cooler Blower Inlet Ducts and
Bearing Airflow Shields have been incorporated. DO NOT incorporate this
Temporary Revision into manual or remove previously issued TR-9, until
modifications 407-799-057 (Inlet Ducts) and 407-799-055 (Bearing Airflow
Shields) or ASB 407-02-54 has been accomplished.
Helicopter S/N 53519 and subsequent will have these modifications
incorporated as basic configuration.
Insert these Temporary Revision pages next to like-number pages in the basic
Flight Manual.
DO NOT remove existing pages. DO NOT remove temporary pages until
replacement pages are received or temporary pages are canceled.
NOTE
For tracking purposes, Temporary Revisions are now being numbered. This
Temporary Revision is issued as TR-10 Revision 1, following previously
issued Temporary Revisions.
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
COPYRIGHT NOTICE
2002
COPYRIGHT
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
RESER
Bell Helicopter
A Subsidiary of Textron Inc.
POST OFFICE BOX 482
FORT WORTH, TEXAS 76101
09 FEBRUARY 1996
TEMPORARY REVISION (TR-10) — 15 FEBRUARY 2002
REVISION 1 —25 JULY 2002
BHT-407-FM-1
LOG OF REVISIONS
Temporary (TR-10).............................15 FEB 02
Revision ...................... 1 .....................25 JUL 02
LOG OF PAGES
PAGE
REVISION
NO.
Title ........................................................... 1
A ................................................................ 1
C-D ............................................................ 1
PAGE
REVISION
NO.
1-7 ......................................... Temporary (TR-10)
2-3 ......................................... Temporary (TR-10)
2-5 — 2-6 .............................. Temporary (TR-10)
NOTICE
CHANGED INFORMATION IN THIS TEMPORARY REVISION APPEARS IN BOLD TYPE.
ALL OTHER INFORMATION IS OUTLINED.
TEMPORARY REVISION (TR-10) — 15 FEB 2002
REVISION 1 — 25 JULY 2002
A
BHT-407-FM-1
LOG OF TC APPROVED REVISIONS
Temporary (TR-10).............................15 FEB 02
Revision ...................... 1 .....................25 JUL 02
DATE
DATE
APPROVED
CHIEF, FLIGHT
FLIGHT TEST
TEST
CHIEF,
FOR
FOR
DIRECTOR —
— AIRCRAFT
AIRCRAFT CERTIFICATION
CERTIFICATION BRANCH
DIRECTOR
DEPARTMENT
OF
TRANSPORT
TRANSPORT CANADA
TEMPORARY REVISION (TR-10) — 15 FEB 2002
REVISION 1 — 25 JULY 2002
C
BHT-407-FM-1
LOG OF FAA APPROVED REVISIONS
Temporary (TR-10) ............................ 22 FEB 02
D
Revision ..................... 1..................... 25 JUL 02
TEMPORARY REVISION (TR-10) — 15 FEB 2002
REVISION 1 — 25 JULY 2002
TC APPROVED
1-13-G.
BHT-407-FM-1
ENGINE OIL TEMPERATURE
Continuous operation
0 to 107°C
Maximum
107°C
If hovering with a tailwind greater than 10
kn o ts at OAT ab o v e 37 . 8° C ( 1 00 ° F ) ,
closely monitor engine oil temperature.
The oil temperature may be reduced by
either turning into wind, reducing power
or transition to forward flight.
NOTE
Positive temperature indication is when
the second segment of the trend arc is
illuminated.
1-14. TRANSMISSION
TEMPORARY REVISION (TR-10) — 15 FEB 2002
1-7
TC APPROVED
BHT-407-FM-1
Section 2
NORMAL PROCEDURES
2-1-B.
HOT WEATHER
OPERATIONS
CAUTION
IF HOV ERIN G W ITH A TAILW IN D
GR EAT ER TH AN 10 K NO TS A T
O AT A B O V E 3 7 . 8 ° C ( 1 0 0 ° F ) ,
CLOSELY MONITOR ENGINE
OIL TEMPERATURE. THE OIL
TEMPERATURE MAY BE
RE DUC ED BY E ITH ER TU RNING
INTO WIND , R EDU CING PO WE R
OR TR AN SITIO N TO FO RWAR D
FLIGH T.
2-2.
FLIGHT PLANNING
TEMPORARY REVISION (TR-10) — 15 FEB 2002
2-3
TC APPROVED
BHT-407-FM-1
f.
Hydromechanical unit —
Security and condition;
evidence of leakage.
g.
Hoses and tubing — Chafing,
security, and condition.
h.
Oil cooler blower inlet duct and
screen — Clear of obstructions,
condition and security.
17. Engine cowl — Secured.
18. Generator cooling scoop —
Clear of debris.
TEMPORARY REVISION (TR-10) — 15 FEB 2002
2-5
BHT-407-FM-1
TC APPROVED
f.
Hydromechanical unit —
Security and condition;
evidence of leakage.
j.
Rotor brake disc and caliper (if
installed) — Condition, security
o f a t t a c h m e n t a n d l e a k a g e.
Ensure brake pads are retracted
from brake disc.
j1.
Oil cooler blower inlet duct and
screen — Clear of obstructions,
condition and security.
k.
Engine cowling — Secured.
l.
Air induction cowling — Secured.
17. Engine cowl — Secured.
18. Generator cooling scoop —
Clear of debris.
2-6
TEMPORARY REVISION (TR-10) — 15 FEB 2002
BHT-407-FM-1
ROTORCRAFT
FLIGHT MANUAL
TEMPORARY REVISION
FOR
FADEC SOFTWARE
VERSION 5.356
(REVERSIONARY GOVERNOR)
Insert these Temporary Revision pages opposite like
numbered pages in the basic Flight Manual after helicopter
is configured with FADEC Software Version 5.356.
DO NOT remove existing pages from Flight Manual. DO NOT
remove temporary pages until replacement pages are
received or temporary pages are canceled.
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
COPYRIGHT NOTICE
COPYRIGHT
2006
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
POST OFFICE BOX 482
FORT WORTH, TEXAS 76101
REISSUE — 17 DECEMBER 2002
TEMPORARY REVISION (TR-11) — 16 NOVEMBER 2006
BHT-407-FM-1
LOG OF REVISIONS
Temporary (TR-11)....0 ..................... 16 NOV 06
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
FLIGHT MANUAL
Title ................................ Temporary (TR-11)
A/B.................................. Temporary (TR-11)
C/D.................................. Temporary (TR-11)
E/F .................................. Temporary (TR-11)
1-3................................... Temporary (TR-11)
1-15................................. Temporary (TR-11)
2-9................................... Temporary (TR-11)
NOTICE
Changed information in this Temporary Revision appears in bold type.
All other information is outlined.
Temporary Revision (TR-11) — 16 NOV 2006———A/B
BHT-407-FM-1
LOG OF TC APPROVED REVISIONS
Temporary (TR-11)....0 ..................... 16 NOV 06
APPROVED
DATE
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
TEMPORARY REVISION (TR-11) — 16 NOV 2006———C/D
BHT-407-FM-1
LOG OF FAA APPROVED REVISIONS
Temporary (TR-11) ... 0 ..................... 16 NOV 06
TEMPORARY REVISION (TR-11) — 16 NOV 2006———E/F
TC APPROVED
BHT-407-FM-1
Section 1
LIMITATIONSTIONS
1
1-5. CONFIGURATION
1-5-A. REQUIRED EQUIPMENT
A functional flashlight is required for night
flights.
FADEC system software shall be version
5.356.
TEMPORARY REVISION (TR-11) — 16 NOV 2006———1-3
TC APPROVED
BHT-407-FM-1
FADEC SOFTWARE VERSION 5.356
WITH DIRECT REVERSION TO
MANUAL INSTALLED, REFER TO
FLIGHT MANUAL FOR OPERATION.
Location: Instrument panel.
407FM_TR11_01
Figure 1-3. Placards and decals (Sheet 3 of 3)
TEMPORARY REVISION (TR-11) — 16 NOV 2006———1-15
TC APPROVED
BHT-407-FM-1
c.
After 3.5 seconds; ENG OUT,
FA D E C D E G R A D E , FA D E C
FAULT, RESTART FAULT, and
ENGINE OVSPD lights illuminate
with activation of engine out
audio for 3 seconds.
d.
Sequence
time).
e.
ENG OUT light re-illumination
with reactivation of engine out
audio after 3 seconds (third
time).
repeats
(second
18. HORN MUTE button — Press to
mute.
19. Caution lights — ENG OUT, XMSN
O I L P R E S S, R P M , H Y D R AU L I C
SYSTEM, GEN FAIL, L/FUEL BOOST,
R/FUEL BOOST, L/FUEL XFR, and
R/FUEL XFR will be illuminated.
TEMPORARY REVISION (TR-11) — 16 NOV 2006———2-9
BHT-407-FM-1
ROTORCRAFT
FLIGHT MANUAL
POST OFFICE BOX 482 •FORT WORTH, TEXAS 76101
BHT-407-FM-1
TEMPORARY REVISION
THIS PAGE HAS
ASSOCIATED
TEMPORARY REVISIONS.
ROTORCRAFT
FLIGHT MANUAL
TR-9
TR-10
TR-11
H-92
TYPE CERTIFICATE NO. ________
REGISTRATION NO. ______________________ SERIAL NO. ____________________
APPROVED BY
DATE _____________________
9 FEBRUARY 1996
DIRECTOR — AIRCRAFT CERTIFICATION BRANCH
DEPARTMENT OF TRANSPORT
THE AVIATION REGULATORY AUTHORITY FOR THIS FLIGHT MANUAL IS THE
CANADIAN DEPARTMENT OF TRANSPORT, AIRCRAFT CERTIFICATION BRANCH.
U.S. REGISTERED HELICOPTERS ARE APPROVED BY THE FAA IN ACCORDANCE
WITH THE PROVISIONS OF 14 CFR SECTION 21.29 ON 23 FEBRUARY 1996
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
COPYRIGHT NOTICE
COPYRIGHT
2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
REISSUE — 17 DECEMBER 2002
REVISION 7 — 30 JUL 2008
BHT-407-FM-1
NOTICE PAGE
The following Warning is not applicable to helicopters on which all kits and customizing
installations have been qualified and approved by Bell Helicopter.
WARNING
THE HELICOPTER MAY CONTAIN INSTALLATIONS, PARTS, OR
PROCESSES CERTIFIED BY PARTIES OTHER THAN BELL
HELICOPTER TEXTRON. BELL HELICOPTER CAN NOT
CONFIRM THAT SUCH INSTALLATIONS HAVE BEEN FULLY
QUALIFIED OR CONFORMED TO BELL HELICOPTER DESIGN
CRITERIA. AS A RESULT OF SUCH INSTALLATIONS, BELL
HELICOPTER
SUPPLIED
DATA
MAY
NOT
BE
VALID
CONCERNING IN-FLIGHT HANDLING QUALITIES, WEIGHT AND
BALANCE, OR HELICOPTER PERFORMANCE. IF MULTIPLE STC
KITS OR SIMILAR INSTALLATIONS ARE INCORPORATED,
THERE MAY BE NOT VALID TEST DATA TO QUALIFY THE
HELICOPTER AS MODIFIED BY THESE INSTALLATIONS. FOR
REVISED DATA, CONTACT THE OWNER OF THE INSTALLED
STC OR THE SUPPLIER FOR THE APPLICABLE APPROVAL OF
EACH INSTALLATION.
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
17 DEC 2002
BHT-407-FM-1
LOG OF REVISIONS
Original ......................0 ......................09 FEB 96
Revision.....................1 .....................08 MAR 96
Revision.....................2 ..................... 09 MAY 96
Revision.....................3 ...................... 30 JUL 96
Revision.....................4 ..................... 04 NOV 96
Revision.....................5 ......................24 JUN 97
Revision.....................6 ...................... 03 JUL 98
Revision.....................7 ......................04 SEP 98
Revision.....................8 ..................... 17 APR 00
Reissue ..................... 0 ......................17 DEC 02
Revision .................... 1 ..................... 10 MAR 04
Revision .................... 2 ......................29 NOV 04
Revision .................... 3 ......................26 APR 05
Revision .................... 4 ...................... 29 JUN 05
Revision .................... 5 ...................... 19 FEB 07
Revision .................... 6 ...................... 20 JUN 07
Revision .................... 7 .......................30 JUL 08
LOG OF PAGES
PAGE
REVISION
NO.
Cover............................................................ 0
Title .............................................................. 7
NP................................................................. 0
A/B................................................................ 7
C/D................................................................ 7
E/F ................................................................ 7
i – ii............................................................... 0
iii/iv............................................................... 0
1-1 – 1-3 ....................................................... 5
1-4 – 1-8 ....................................................... 7
1-9 – 1-13 ..................................................... 0
1-14............................................................... 5
1-15............................................................... 7
1-16 – 1-17 ................................................... 0
1-18 – 1-19 ................................................... 7
1-20............................................................... 3
2-1/2-2 .......................................................... 6
2-3 – 2-7 ....................................................... 5
2-8 – 2-9 ....................................................... 7
2-10 – 2-11 ................................................... 5
PAGE
REVISION
NO.
2-12 – 2-15 ................................................... 6
2-16 .............................................................. 7
2-17/2-18...................................................... 0
3-1/3-2 .......................................................... 7
3-3 ................................................................ 7
3-4 – 3-8 ....................................................... 0
3-9 ................................................................ 7
3-10 – 3-12 ................................................... 0
3-13 .............................................................. 5
3-14 – 3-17 ................................................... 0
3-18 – 3-19 ................................................... 5
3-20 .............................................................. 0
4-1/4-2 .......................................................... 0
4-3 – 4-50 ..................................................... 0
4-51/4-52...................................................... 0
5-1 – 5-18 ..................................................... 0
5-19/5-20...................................................... 0
A-1/A-2......................................................... 5
A-3 – A-4...................................................... 5
A-5/A-6......................................................... 5
TEMPORARY REVISION
THIS PAGE HAS
ASSOCIATED
TEMPORARY REVISIONS.
TR-9
TR-10
TR-11
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
30 JUL 2008—Rev. 7———A/B
BHT-407-FM-1
LOG OF TC APPROVED REVISIONS
Original ..................... 0...................... 09 FEB 96
Revision .................... 1..................... 08 MAR 96
Revision .................... 2......................09 MAY 96
Revision .................... 3....................... 30 JUL 96
Revision .................... 4...................... 04 NOV 96
Revision .................... 5...................... 24 JUN 97
Revision .................... 6....................... 03 JUL 98
Revision .................... 7...................... 04 SEP 98
Revision .................... 8...................... 17 APR 00
Reissue...................... 0 ..................... 17 DEC 02
Revision .................... 1 ..................... 10 MAR 04
Revision .................... 2 ..................... 29 NOV 04
Revision .................... 3 ..................... 26 APR 05
Revision .................... 4 ...................... 29 JUN 05
Revision .................... 5 ...................... 19 FEB 07
Revision .................... 6 ...................... 20 JUN 07
Revision .................... 7 ...................... 30 JUL 08
TEMPORARY REVISION
THIS PAGE HAS
ASSOCIATED
TEMPORARY REVISIONS.
TR-9
TR-10
TR-11
APPROVED
DATE
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
30 JUL 2008—Rev. 7———C/D
BHT-407-FM-1
LOG OF FAA APPROVED REVISIONS
Original ..................... 0...................... 09 FEB 96
Revision .................... 1..................... 08 MAR 96
Revision .................... 2......................09 MAY 96
Revision .................... 3....................... 30 JUL 96
Revision .................... 4...................... 04 NOV 96
Revision .................... 5...................... 24 JUN 97
Revision .................... 6......................18 MAY 99
Revision .................... 7......................18 MAY 99
Revision .................... 8......................05 MAY 00
Reissue...................... 0 ...................... 24 SEP 03
Revision .................... 1 ..................... 11 MAR 04
Revision .................... 2 ..................... 07 DEC 04
Revision .................... 3 ..................... 04 MAY 05
Revision .................... 4 ...................... 30 JUN 05
Revision .................... 5 ..................... 08 MAR 07
Revision .................... 6 ..................... 31 AUG 07
Revision .................... 7 ..................... 18 AUG 08
TEMPORARY REVISION
THIS PAGE HAS
ASSOCIATED
TEMPORARY REVISIONS.
TR-9
TR-10
TR-11
30 JUL 2008—Rev. 7———E/F
BHT-407-FM-1
GENERAL INFORMATION
ORGANIZATION
TERMINOLOGY
This Rotorcraft Flight Manual is divided into
five sections and an appendix as follows:
WARNINGS, CAUTIONS, AND
NOTES
Section 1
Section 2
Section 3
Section 4
Section 5
Appendix A
— LIMITATIONS
— NORMAL PROCEDURES
— EMERGENCY AND
MALFUNCTION
PROCEDURES
— PERFORMANCE
— WEIGHT AND BALANCE
— OPTIONAL EQUIPMENT
SUPPLEMENTS
Sections 1 through 4 contain Transport
Canada (TC) approved data necessary to
operate basic helicopter in a safe and
efficient manner.
Section 5 contains weight and balance data
necessary for flight planning.
Appendix A contains a list of approved
supplements for optional equipment, which
shall be used in conjunction with basic
Flight Manual when respective optional
equipment kits are installed.
Warnings, cautions, and notes are used
throughout this manual to emphasize
important and critical instructions as
follows:
WARNING
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
Manufacturer's Data manual (BHT-407-MD-1)
contains information to be used in
conjunction
with
Flight
Manual.
Manufacturer's data manual is divided into
four sections:
Section 1
— SYSTEMS DESCRIPTION
Section 2
— HANDLING AND
SERVICING
Section 3
— CONVERSION CHARTS
AND TABLES
Section 4
— EXPANDED
PERFORMANCE
An operating procedure, condition,
etc., which is essential to highlight.
USE OF PROCEDURAL WORDS
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 when
application of a procedure is recommended.
17 DEC 2002
i
BHT-407-FM-1
MAY and NEED NOT have been used only
when application of a procedure is optional.
dBA
— Decibel, “A” type filter
DC
— Direct current
WILL has been used only to indicate
futurity, never to indicate a mandatory
procedure.
DG
— Directional gyro
DOT
— Department of Transport
ECS
— Environmental control
system
ECU
— Engine control unit
ELT
— Emergency locator
transmitter
ENCDG
— Encoding
ABBREVIATIONS,
PLACARDING
ACRONYMS
AND
Abbreviations, acronyms and placarding
used throughout this manual are defined as
follows:
ADF
— Automatic direction finder
ENG
— Engine
AIR
COND
— Air conditioner
ENG
ANTI ICE
— Engine anti icing
A/F
— Airframe
°F
— Degrees Fahrenheit
ALT
— Altimeter
FADEC
ANTI
COLL LT
— Anticollision light
— Full authority digital engine
control
FS
— Fuselage station
ATT
— Attitude
FT or ft
— Foot, feet
AUTO
— Automatic
FWD
— Forward
AUX
— Auxiliary
GEN
— Generator
BATT
— Battery
GOV
— Governor
BIT
— Built in test
GPS
— Global positioning system
BL
— Buttock line
GPU
— Ground power unit
BLO
— Blower
GW
— Gross weight
BRT
— Bright
HD
— Density altitude
°C
— Degrees Celsius
HG
— Inches of mercury
CAUT
— Caution
HMU
— Hydromechanical unit
CAUT LT
— Caution lights
Hp
— Pressure altitude
CG
— Center of gravity
HYD
— Hydraulic
CKPT
— Cockpit
HV
— Height-velocity
CM
— Centimeter (s)
ICAO
COMM
— Communication
— International Civil Aviation
Organization
CONT
— Control
ICS
— Intercommunication system
IFL
— Inflate
ii
17 DEC 2002
BHT-407-FM-1
IGE
— In ground effect
IGNTR
— Ignitor
IN
— Inch(es)
INSTR
CHK
— Instrument check
INSTR LT
PMA
— Permanent Magnetic
Alternator
POS LT
— Position light
PRESS
— Pressure
PSI
— Pounds per square inch
— Instrument light
PTT
— Press to Test
KCAS
— Knots calibrated airspeed
PWR
— Power
KG or kg
— Kilogram(s)
QTY
— Quantity
KIAS
— Knots indicated airspeed
R/FUEL
— Right fuel
KTAS
— Knots true airspeed
RECP
— Receptacle
L
— Liter(s)
RLY
— Relay
LB(S) or
lb(s)
— Pound(s)
RPM
— Revolutions per minute
RTR
— Rotor
LDG LTS
— Landing lights
L/FUEL
— Left fuel
LT
— Light
MAN
— Manual
MCP
— Maximum continuous
power
MD
— Manufacturer's Data
MGT
— Measured gas temperature
MM or
mm
— Millimeter(s)
NAV
s/w Ver
Soft ware version
SEL
— Sound exposure level
SHP
— Shaft horsepower
SL
— Sea level
SPKR
— Speaker
Sq
— Square
SYS
— System
T/R
— Tail rotor
TCA
— Transport Canada Aviation
— Navigation
TEMP
— Temperature
NG
— Gas producer RPM
TRQ
— Torque
NP
— Power turbine RPM
VFR
— Visual flight rules
NR
— Rotor RPM
VHF
— Very high frequency
OAT
— Outside air temperature
VNE
— Never exceed velocity
OBS
— Omni bearing selector
VOR
— VHF omnidirectional range
OGE
— Out of ground effect
WL
— Water line
OVSPD
— Overspeed
WARN
— Warning
PART
SEP
— Particle separator
XFR
— Transfer
XMSN
— Transmission
PASS
— Passenger(s)
XPDR
— Transponder
17 DEC 2002
iii/iv
TC APPROVED
BHT-407-FM-1
Section 1
LIMITATIONS
1
TABLE OF CONTENTS
Subject
Paragraph
Number
Page
Number
Introduction ............................................................................................
Basis of Certification .............................................................................
Types of Operation ................................................................................
Passengers.........................................................................................
Cargo...................................................................................................
Flight Crew .............................................................................................
Configuration .........................................................................................
Required Equipment..........................................................................
Optional Equipment..........................................................................
Doors Removed ................................................................................
Weight and Center of Gravity ...............................................................
Weight .................................................................................................
Center of Gravity................................................................................
Airspeed..................................................................................................
Altitude....................................................................................................
Maneuvering...........................................................................................
Prohibited Maneuvers .......................................................................
Climb And Descent ............................................................................
Slope Landing ....................................................................................
Not Used .................................................................................................
Ambient Temperatures..........................................................................
Electrical .................................................................................................
Generator............................................................................................
Starter .................................................................................................
Power Plant ............................................................................................
Gas Producer RPM (NG) ....................................................................
Power Turbine RPM (NP) ...................................................................
Measured Gas Temperature (MGT) ..................................................
Engine Torque....................................................................................
Fuel Pressure .....................................................................................
Engine Oil Pressure...........................................................................
Engine Oil Temperature ....................................................................
Transmission..........................................................................................
Transmission Oil Pressure ...............................................................
Transmission Oil Temperature .........................................................
Rotor .......................................................................................................
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 .....
1-13 .........
1-13-A .....
1-13-B .....
1-13-C .....
1-13-D .....
1-13-E .....
1-13-F......
1-13-G .....
1-14 .........
1-14-A .....
1-14-B .....
1-15 .........
1-3
1-3
1-3
1-3
1-3
1-3
1-3
1-3
1-4
1-4
1-4
1-4
1-4
1-4
1-5
1-5
1-5
1-5
1-5
1-5
1-5
1-5
1-5
1-5
1-5
1-6
1-6
1-6
1-6
1-6
1-6
1-7
1-7
1-7
1-7
1-7
19 FEB 2007—Rev. 5———1-1
BHT-407-FM-1
TC APPROVED
TABLE OF CONTENTS (CONT)
Subject
Paragraph
Number
Page
Number
Rotor RPM — Power ON ...................................................................
Rotor RPM — Power OFF..................................................................
Hydraulic ................................................................................................
Fuel and Oil ............................................................................................
Fuel .....................................................................................................
Oil ........................................................................................................
Rotor Brake ............................................................................................
Not Used .................................................................................................
Instrument Markings and Placards ......................................................
1-15-A .....
1-15-B .....
1-16.........
1-17.........
1-17-A .....
1-17-B .....
1-18.........
1-19.........
1-20.........
1-7
1-7
1-7
1-7
1-7
1-8
1-8
1-8
1-8
Figure
Number
1-1...........
1-2...........
1-3...........
1-4...........
1-5...........
Page
Number
1-9
1-11
1-13
1-16
1-17
LIST OF FIGURES
Subject
Gross Weight Longitudinal Center of Gravity Limits .........................
Gross Weight Lateral Center of Gravity Limits...................................
Placards and Decals..............................................................................
Ambient Air Temperature Limitations .................................................
Instrument Markings .............................................................................
1-2———Rev. 5—19 FEB 2007
TEMPORARY REVISION
TC APPROVED
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
Section 1
BHT-407-FM-1
TR-11
LIMITATIONS
1
1-1. INTRODUCTION
Complianc e with Limitations section is
required by appropriate operating rules.
Anytime an operating limitation is exceeded,
an appropriate entry shall be made in
helicopter logbook. Entry shall state which
limit was exceeded, duration of time, extreme
value attained, and any additional information
essential in determining maintenance action
required.
I n t e n t i o n a l u s e o f t r a n s i e n t l i m i ts i s
prohibited.
Torque events shall be recorded. A torque
event is defined as a takeoff or lift, internal or
external load (BHT-407-MD-1).
Landings shall be recorded. Run-on landings
shall be recorded separately.
A run-on landing is defined as one where
there is forward ground travel of the
helicopter greater than 3 feet with the weight
on the skids.
1-2. BASIS OF CERTIFICATION
This helicopter is certified under FARs Parts
27 and 36, Appendix J. Additionally, it is
approved under Canadian Airworthiness
Manual Chapters 516 (ICAO Chapter 11) and
527, Sections 1093 (b) (1) (ii) and (iii), 1301-1,
1557 (c) (3), 1581 (e) and 1583 (h).
1-3. TYPES OF OPERATION
1-3-A.
PASSENGERS
Basic configured helicopter is approved for
seven place seating and is certified for land
operation under day or night VFR non-icing
conditions.
1-3-B.
CARGO
The maximum allowable cabin deck loading
for cargo is 75 pounds per square foot (3.7 kg
per 100 cm2). The maximum allowable
baggage compartment deck loading is 86
pounds per square foot (4.2 kg per 100 cm2)
with a maximum allowable weight of 250
pounds (113.4 kg). Refer to BHT-407-MD-1 for
cargo restraint and tie-down locations.
Cargo must be properly secured by tie-down
devices to prevent the load from shifting
un der a nticipate d fligh t a nd g rou nd
operations. If the mission requires both
passengers and cargo to be transported
together, the cargo must be loaded and
secured so that it does not obstruct
passenger access to exits.
1-4. FLIGHT CREW
Minimum flight crew consists of one pilot who
shall operate helicopter from right crew seat.
Left crew seat may be used for an additional
pilot when approved dual controls are
installed.
1-5. CONFIGURATION
1-5-A.
REQUIRED EQUIPMENT
A functional flashlight is required for night
flights.
FADEC system software shall be version
5.202.
19 FEB 2007—Rev. 5———1-3
BHT-407-FM-1
1-5-B.
TC APPROVED
OPTIONAL EQUIPMENT
CAUTION
The snow deflector kit (BHT-407-FMS-4) shall
b e i n s ta l l e d w h e n c o n d u c t i n g f l i g h t
operations in falling and/or blowing snow.
Refer to appropriate flight manual
supplement(s) (FMS) for additional
limitations, procedures, and performance data
for optional equipment.
1-5-C.
DOORS REMOVED
NOTE
Indicated altitude may be up to 100
feet lower than actual altitude with
crew door(s) removed.
Flight with any combination of doors removed
is approved. With litter door removed, left
passenger door shall be removed. Refer to
Airspeed limitations.
With door(s) removed, determine weight
change and adjust ballast if necessary. Refer
to Section 5.
NOTE
All unsecured items shall be
removed from cabin when any door
is removed.
1-6. WEIGHT AND CENTER OF
GRAVITY
1-6-A.
LOADS THAT RESULT IN GW
ABOVE 5000 POUNDS (2268 KG)
SHALL BE CARRIED ON THE
CARGO HOOK AND MUST BE
JETTISONABLE.
Maximum approved GW for flight with
jettisonable external load is 6000 pounds
(2722 kg).
1-6-B.
CENTER OF GRAVITY
The pilot is responsible for determining
weight and balance to ensure gross weight
and center of gravity will remain within limits
throughout each flight. Refer to Section 5 for
loading tables and instructions.
NOTE
Ballast as required to maintain most
forward or most aft CG within GW
flight limits (Figure 1-1). For standard
passenger and fuel loadings,
applicable Weight Empty Center of
Gravity Chart in BHT-407-MM-1 may
b e u s e d t o d e t e r m in e r e q u i r e d
ballast.
For longitudinal CG limits, refer to Gross
Weight Longitudinal Center of Gravity Limits
chart (Figure 1-1).
For lateral CG limits, refer to Gross Weight
Lateral Center of Gravity Limits (Figure 1-2).
WEIGHT
Maximum approved internal GW for takeoff
and landing is 5000 pounds (2268 kg).
1-7. AIRSPEED
M i n im u m G W fo r fl ig h t is 2 6 5 0 p o u n d s
(1202 kg).
Basic VNE is 140 KIAS, sea level to 3000 feet
H D. Decrease V NE for ambient conditions in
accordance with AIRSPEED LIMITATIONS
Placards and Decals (Figure 1-3).
Minimum weight at fuselage station 65.0 is
170 pounds (77.1 kg).
VNE at 93.5 to 100% TORQUE (takeoff power)
is 100 KIAS, not to exceed placarded VNE.
1-4———Rev. 7—30 JUL 2008
TR-9
TEMPORARY REVISION
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
TC APPROVED
BHT-407-FM-1
VNE is 100 KIAS or placarded VNE, whichever
is less, when takeoff loading is in shaded area
of the Gross Weight Lateral Center of Gravity
Limits (Figure 1-2).
Slope landings are limited to 10° side slopes,
10° nose up slope or 5° nose down slope.
VNE is 100 KIAS with any door(s) removed, not
to exceed placarded VNE.
1-11. AMBIENT TEMPERATURES
VNE is 100 KIAS or placarded VNE, whichever
is less for steady state autorotation.
Maximum allowable airspeed for sideward
and rearward flight or crosswind hover is 35
KTAS.
1-8. ALTITUDE
1-10. NOT USED
Maximum sea level ambient air temperature
for operation is 51.7°C (125°F) and decreases
with H P at standard lapse rate of 2°C (3.6°F)
per 1000 feet. Refer to Ambient Air
Temperature Limitations chart (Figure 1-4).
Minimum ambient air temperature for
operation at all altitudes is -40°C (-40°F).
ENG ANTI ICE shall be ON in visible moisture
when OAT is below 5°C (40°F).
Maximum operating altitude is 20,000 feet HD
or 20,000 feet HP, whichever is lower.
1-12. ELECTRICAL
1-9. MANEUVERING
1-12-A. GENERATOR
1-9-A.
PROHIBITED MANEUVERS
Aerobatic maneuvers are prohibited.
1-9-B.
CLIMB AND DESCENT
Maximum rate o f c limb is 2000 feet per
minute.
1-9-C.
SLOPE LANDING
CAUTION
SLOPE LANDINGS HAVE BEEN
DEMONSTRATED TO THE SLOPE
LANDING
LIMITS.
OTHER
CONDITIONS INCLUDING, BUT NOT
LIMITED TO, WIND DIRECTION AND
VELOCITY, CENTER OF GRAVITY,
AND THE CONDITION OF THE
SLOPE (LOOSE ROCK, SOFT MUD,
SNOW, WET GRASS, ETC.) MAY
LIMIT MAXIMUM SLOPE TO A VALUE
LESS THAN THE PUBLISHED
LIMITS.
Continuous operation,
up to 10,000 feet Hp
0 to 180 amps
Maximum continuous up
to 10,000 feet Hp
180 amps
Continuous operation,
above 10,000 feet Hp
0 to 170 amps
Maximum continuous
above 10,000 feet Hp
170 amps
Transient, 2 minutes
180 to 300 amps
Transient, 5 seconds
300 to 400 amps
1-12-B. STARTER
External Power Start
Battery Start
40 seconds ON
60 seconds ON
30 seconds OFF
60 seconds OFF
40 seconds ON
60 seconds ON
30 seconds OFF
60 seconds OFF
40 seconds ON
60 seconds ON
30 minutes OFF
30 minutes OFF
30 JUL 2008—Rev. 7———1-5
BHT-407-FM-1
TC APPROVED
NOTE
NOTE
28 VDC GPU for starting shall be
limited to 500 amps.
1-13. POWER PLANT
Rolls-Royce model 250-C47B.
NOTE
Intentional use of any power
transient is prohibited.
1-13-A. GAS PRODUCER RPM (NG)
NOTE
FADEC will limit N G in accordance
with GAS PRODUCER RPM (N G )
LIMIT placard. NR decay will result if
power demand exceeds placard limit.
Maximum continuous NG is limited in
accordance with GAS PRODUCER RPM (N G)
LIMIT placard (Figure 1-3) when operating
above 10,000 feet H P and with OAT below
-30°C (-22°F).
Continuous operation
63 to 105%
Maximum continuous
operation
105%
Transient, 10 seconds
105.1 to 106%
1-13-B. POWER TURBINE RPM (NP)
Avoid continuous
operations
68.4 to 87.1%
Minimum
99%
Continuous operation
99 to 100%
Maximum continuous
100%
Maximum transient, 15
seconds
102.1 to 107% NP
1-6———Rev. 7—30 JUL 2008
ENGINE OVSPD warning light will
illuminate when NP versus TORQUE
is between 102.4% NP at 100%
TORQUE and 108.6% NP at 0%
TORQUE.
When operating in MANUAL mode NP
should be maintained between 95
and 100%.
1-13-C. MEASURED GAS TEMPERATURE
(MGT)
GAUGE P/N 407-375-001-101/-103
Continuous operation
100 to 727°C
Maximum continuous
727°C
Takeoff, 5 minutes
727 to 779°C
Maximum for takeoff
779°C
Transient, 12 seconds
780 to 826°C
Maximum starting, do not
exceed 10 seconds above
826°C or 1 second at
927°C.
927°C
NOTE
Either MGT gauge may be installed.
GAUGE P/N 407-375-001-105 AND SUB
Continuous operation
100 to 727°C
Maximum continuous
727°C
Takeoff, 5 minutes
727 to 779°C
Maximum for takeoff
779°C
Transient, 12 seconds
780 to 905°C
Maximum starting, do not
exceed 10 seconds above
843°C or 1 second at
927°C.
927°C
TC APPROVED
TR-9
TEMPORARY REVISION
TR-10
THIS PAGE HAS
ASSOCIATED
TEMPORARY REVISIONS.
BHT-407-FM-1
1-14. TRANSMISSION
1-13-D. ENGINE TORQUE
Continuous operation
0 to 93.5%
Maximum continuous
93.5%
Takeoff, 5 minute
93.5 to 100%
Transient, 5 seconds
105%
NOTE
Use of takeoff power is limited to 100
KIAS, not to exceed placarded VNE.
1-13-E. FUEL PRESSURE
Minimum
8 PSI
Continuous operation
8 to 25 PSI
Maximum
25 PSI
1-13-F. ENGINE OIL PRESSURE
Minimum below 79% NG
50 PSI
Minimum from 79 to 94%
NG
90 PSI
Minimum above 94% NG
115 PSI
Maximum
130 PSI
Maximum cold starts only 200 PSI
NOTE
When 130 PSI is exceeded during
start, operate engine at idle until oil
pressure drops below 130 PSI.
1-13-G. ENGINE OIL TEMPERATURE
Continuous operation
0 to 107°C
Maximum
107°C
NOTE
Positive temperature indication is
when the second segment of the
trend arc is illuminated.
1-14-A. TRANSMISSION OIL PRESSURE
Minimum
30 PSI
Continuous operation
40 to 70 PSI
Maximum
70 PSI
1-14-B. TRANSMISSION OIL
TEMPERATURE
Continuous operation
15 to 110°C
Maximum
110°C
1-15. ROTOR
1-15-A. ROTOR RPM — POWER ON
Continuous operation
99 to 100%
Maximum continuous
100%
NOTE
When operating in MANUAL mode
N R should be maintained between
95% and 100%.
1-15-B. ROTOR RPM — POWER OFF
Minimum
85%
Continuous operation
85 to 107%
Maximum
107%
CAUTION
FOR AUTOROTATIVE TRAINING,
MAINTAIN STEADY STATE NR
ABOVE 90%.
1-16. HYDRAULIC
Hydraulic fluid MIL-PRF-5606 (NATO H-515)
may be used at all ambient temperatures.
30 JUL 2008—Rev. 7———1-7
BHT-407-FM-1
TC APPROVED
1-17. FUEL AND OIL
1-17-B-2.
OIL — TRANSMISSION AND TAIL
ROTOR GEARBOX
1-17-A. FUEL
NOTE
Fuel conforming to following specifications
may be used at all ambient temperatures:
ASTM-D-6615, Jet B
MIL-DTL-5624, Grade JP-4 (NATO F-40)
Fuels conforming to following specifications
are limited to ambient temperatures of -32°C
(-25°F) and above:
ASTM-D-1655, Jet A or A-1
MIL-DTL-5624, Grade JP-5 (NATO F-44)
MIL-DTL-83133, Grade JP-8 (NATO F-34).
For operations below -32°C (-25°F), refer to
Rolls-Royce Operation and Maintenance
Manual for cold weather fuel and blending
instructions.
It is recommended DOD-PRF-85734
oil be used in transmission and tail
rotor gearbox to maximum extent
allowed by temperature limitations.
Oil conforming to DOD-PRF-85734 is limited
to ambient temperatures above -40°C (-40°F).
Oil conforming to MIL-PRF-7808 (NATO O-148)
is limited to ambient temperatures below
-18°C (0°F).
1-18. ROTOR BRAKE
Rotor brake (if installed) application is limited
to ground operation after engine has been
shut down and N R has decreased to 40% or
lower.
For emergency stops, apply rotor brake any
time after engine is shut down.
1-17-B. OIL
Engine starts with rotor brake engaged are
prohibited.
1-17-B-1.
1-19. NOT USED
OIL — ENGINE
Oil con form ing to MIL-PRF -7 808 (NATO
O-148), DOD-PRF-85734 or MIL-PRF-23699
( N AT O O - 1 5 6 ) i s l i m i t e d t o a m b i e n t
temperatures above -40°C (-40°F).
NOTE
Refer to Rolls-Royce Operation and
Maintenance
Manual
and
BHT-407-MD-1 manual for approved
oils and mixing of oils of different
brands, types, and manufacturers.
1-8———Rev. 7—30 JUL 2008
1-20. INSTRUMENT MARKINGS
AND PLACARDS
Refer to Figure 1-3 for Placards and Decals.
Refer to Figure 1-5 for Instrument Markings.
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.
TC APPROVED
BHT-407-FM-1
LONGITUDINAL C.G.
6200
6000
120.5
127.6
6000
5800
5600
EXTERNAL LOAD ONLY
5400
5200
GROSS WEIGHT - POUNDS
5000
5000
119.5
4800
4600
4500
4400
4200
4000
3800
3600
3400
3200
3000
2800
2600
2400
118
2800
119.0
119
2650
120
121
122
123
124
125
126
127
128.0
129.0
128
129
130
FUSELAGE STATION - INCHES
M407_FM-1__FIG_1-1_(1_OF_2).WMF
Figure 1-1. Gross weight longitudinal center of gravity limits (Sheet 1 of 2)
17 DEC 2002
1-9
BHT-407-FM-1
TC APPROVED
LONGITUDINAL C.G.
2800
3061
3241
2722
2700
2600
2500
EXTERNAL LOAD ONLY
2400
2300
GROSS WEIGHT - KILOGRAMS
2268
3035
2200
2100
2041
2000
1900
1800
1700
1600
1500
1400
1300
1270
3023
1200
1100
3000
3025
1202
3050
3075
3100
3125
3150
3175
3200
3225
3251
3277
3250
3275
3300
FUSELAGE STATION - MILLIMETERS
M407_FM-1__FIG_1-1_(2_OF_2).WMF
Figure 1-1. Gross weight longitudinal center of gravity limits (Sheet 2 of 2)
1-10
17 DEC 2002
TC APPROVED
BHT-407-FM-1
LATERAL C.G.
6200
-0.9
6000
1.4
6000
5800
5600
EXTERNAL LOAD ONLY
5400
MAX AIRSPEED 100 KIAS
5200
GROSS WEIGHT - POUNDS
5000
5000
MAX AIRSPEED
100 KIAS
4800
4600
MAX AIRSPEED
100 KIAS
-1.5
EXTERNAL
LOAD ONLY
2.0
EXTERNAL
LOAD ONLY
4400
4200
4000
3800
3600
3500
3500
3400
3200
3000
2800
2650
2600
-4.0
-2.5
3.0
4.0
3
4
2400
-5
-4
-3
-2
-1
0
1
2
5
BUTTOCK LINE - INCHES
M407_FM-1__FIG_1-2_(1_OF_2).WMF
Figure 1-2. Gross weight lateral center of gravity limits (Sheet 1 of 2)
17 DEC 2002
1-11
BHT-407-FM-1
TC APPROVED
LATERAL C.G.
2800
-23
36
2722
2700
2600
2500
EXTERNAL LOAD ONLY
MAX AIRSPEED 100 KIAS
2400
2300
GROSS WEIGHT - KILOGRAMS
2268
2200
MAX AIRSPEED
100 KIAS
2100
EXTERNAL
LOAD ONLY
MAX AIRSPEED
100 KIAS
52
-39
EXTERNAL
LOAD ONLY
2000
1900
1800
1700
1600
1588
1588
1500
1400
1300
1202
1200
-102
1100
-125
-100
-64
-75
-50
-25
0
25
50
76
102
75
100
125
BUTTOCK LINE - MILLIMETERS
M407_FM-1__FIG_1-2_(2_OF_2).WMF
Figure 1-2. Gross weight lateral center of gravity limits (Sheet 2 of 2)
1-12
17 DEC 2002
TC APPROVED
BHT-407-FM-1
EMERGENCY PEDAL
STOP RELEASE
PULL ONLY
MAINT. RESET
REQUIRED
Location: Between Pilot and Copilot seats
407 AIRSPEED LIMITATIONS - KIAS
OAT
°C
52
45
40
35
30
25
20
0
-25
-40
PRESSURE ALTITUDE FT x 1000
0
137
139
140
140
140
140
140
140
140
137
2
4
6
8
10
12
14
16
18
20
132
133
135
137
138
140
140
140
133
125
126
128
129
131
133
140
140
128
119
120
122
124
125
132
135
123
113
115
116
118
125
130
118
108
109
111
117
125
114
102
103
110
119
110
95
96
103
111
105
89
95
104
101
88
97
97
89
93
MAXIMUM AUTOROTATION VNE 100 KIAS
Airspeed limits shown are valid only for corresponding altitudes and temperatures.
Hatched areas indicate conditions which exceed approved temperature or density altitude
limitations.
Location: Forward of Overhead Console
M407_FM-1__FIG_1-3_(VNE).WMF
Figure 1-3. Placards and decals (Sheet 1 of 3)
17 DEC 2002
1-13
BHT-407-FM-1
TC APPROVED
FUEL
FUEL SYSTEM USABLE CAPACITY
BASIC AIRCRAFT 127 U.S. GALLONS - 483 LITERS
WITH 407-706-011 AUX KIT 147 U.S. GALLONS = 559 LITERS
SEE FLIGHT MANUAL FOR APPROVED FUELS
Location: Above fuel filler cap.
AVOID CONT OPS 68.4% TO 87.1% NP
Location: Instrument panel.
THIS HELICOPTER MUST BE OPERATED IN
COMPLIANCE WITH THE OPERATING LIMITATIONS
SPECIFIED IN THE APPROVED FLIGHT MANUAL
Location: Bottom and centered on instrument panel.
DO NOT APPLY ROTOR BRAKE
ABOVE 40% RPM
Location: Near rotor brake (if installed).
CARGO MUST BE SECURED
IN ACCORDANCE WITH
FLIGHT MANUAL INSTR
Location: Inside of baggage door.
407-FM-1-3-2
Figure 1-3. Placards and decals (Sheet 2 of 3)
1-14
Rev. 5
19 FEB 2007
TC APPROVED
BHT-407-FM-1
TEMPORARY REVISION
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
TR-11
GAS PRODUCER RPM (NG) LIMIT
WHEN ABOVE 10,000 FT HP
MAXIMUM Ng % RPM WITH OAT IS AS FOLLOWS
OAT
ºC
-40
-39
-38
-37
-36
-35
-34
-33
-32
-31
-30
MAX 99.0 99.2 99.4 99.6 99.8 100.0 100.2 100.4 100.6 100.8 101.1
Ng %
Location: Above pilot windshield
FADEC SOFTWARE VERSION 5.202
WITH DIRECT REVERSION TO
MANUAL INSTALLED. REFER TO
FLIGHT MANUAL FOR OPERATION
Location: Instrument panel
MAX ALLOWABLE WEIGHT 250 LBS.
MAX ALLOWABLE WEIGHT PER SQ. FT. 86 LBS.
Location: Inside of baggage door
FUEL CAPACITY
BASIC 869 LBS
WITH AUX 1005 LBS
(JET A AT 15ºC)
Location: Instrument panel
Location: Instrument panel and passenger compartment
407_FM_1_0002
Figure 1-3. Placards and Decals (Sheet 3 of 3)
30 JUL 2008
Rev. 7
1-15
BHT-407-FM-1
TC APPROVED
AMBIENT AIR TEMPERATURE - °F
-40
-20
0
20,000
20
MA
X IM
UM
40
H
D
18,000
FL
IG
60
HT
L IM
80
100
120
140
IT
AT
8,000
MO
10,000
IMU
12,000
MINIMUM OAT FLIGHT LIMIT
14,000
X
MA
GH
F LI
PRESSURE ALTITUDE - FT
16,000
6,000
T
IM I
TL
EXAMPLE: AT HP = 7000 FT
MAX OAT = 37.8°C (100°F)
MIN OAT = - 40°C (- 40°F)
4,000
2,000
0
-40
-30
-20
-10
0
10
20
30
40
50
60
AMBIENT AIR TEMPERATURE - °C
M407_FM-1__FIG_1-4.EMF
Figure 1-4. Ambient air temperature limitations
1-16
17 DEC 2002
TC APPROVED
BHT-407-FM-1
ENGINE OIL PRESSURE
50 PSI
Minimum
50 to 90 PSI
Operation below 79% NG RPM
90 to 115 PSI
Continuous operation below 94% NG RPM
115 to 130 PSI
Continuous operation
130 PSI
Maximum for continuous operation
200 PSI
Maximum for cold start
ENGINE OIL TEMPERATURE
0 to 107°C
Continuous operation
107°C
Maximum
TRANSMISSION OIL PRESSURE
30 PSI
Minimum
40 to 70 PSI
Continuous operation
70 PSI
Maximum
TRANSMISSION OIL TEMPERATURE
15 to 110°C
Continuous operation
110°C
Maximum
NG (GAS PRODUCER RPM)
63 to 105%
Continuous operation
105%
Maximum continuous operation
106%
Maximum transient, 10 seconds
M407_FM-1__FIG_1-5_(1_OF_4).EPS
Figure 1-5. Instrument markings (Sheet 1 of 4)
17 DEC 2002
1-17
BHT-407-FM-1
TC APPROVED
TRQ (TORQUE)
0 to 93.5%
Continuous operation
93.5 to 100%
5 minute takeoff range
100%
Maximum
407-375-001-105 AND SUB.
MGT (MEASURED GAS TEMPERATURE)
*
100 to 727°C
Continuous operation
727 to 779°C
5 minute takeoff range
779°C
Maximum for takeoff
826°C
or
843°C
Beginning of 10 seconds
range for starting
927°C
Maximum for start and shutdown
(1 second maximum)
* Either gauge may be installed
AIRSPEED
0 to 140 Knots
Continuous operation
100 Knots
Maximum for autorotation
140 Knots
Maximum
407_FM_1_0003
Figure 1-5. Instrument Markings (Sheet 2 of 4)
1-18
Rev. 7
30 JUL 2008
TC APPROVED
BHT-407-FM-1
407-375-008-101/-103
NP (POWER TURBINE RPM)
99%
Minimum
99 to 100%
Continuous operation
100%
Maximum continuous
NR (ROTOR RPM)
85%
Minimum (power off)
85 to 107%
Continuous operation (power off)
107%
Maximum (power off)
FUEL QUANTITY
(Jet A 6.8 lbs/gal)
0 LBS
All tanks empty (zero useable)
195 LBS
Forward tank empty
869 LBS
Forward and aft tanks full
1005 LBS
Forward, aft and auxiliary
tanks full
407_FM_1_0004
Figure 1-5. Instrument Markings (Sheet 3 of 4)
30 JUL 2008
Rev. 7
1-19
BHT-407-FM-1
TC APPROVED
* Either gauge may be installed.
* P/N 407-075-024-101
DC LOAD
170 Amps
Maximum continuous above 10,000 FT Hp
180 Amps
Maximum
FUEL PRESSURE
8 PSI
Minimum
8 to 25 PSI
Continuous operation
25 PSI
Maximum
407-075-024-103,
* P/N
407-375-007-105 or 407-375-007-107
DC LOAD
0 to 180 Amps
Continuous operation
170 Amps
Maximum continuous above 10,000 FT Hp
180 Amps
Maximum
300 Amps
Maximum transient, 2 minutes
400 Amps
Maximum transient, 5 seconds
FUEL PRESSURE
8 PSI
Minimum
8 to 25 PSI
Continuous operation
25 PSI
Maximum
VERTICAL SPEED INDICATOR
2,000 Feet per minute up
Maximum
407FM_1_0001
Figure 1-5. Instrument markings (Sheet 4 of 4)
1-20
Rev. 3 26 APR 2005
TC APPROVED
BHT-407-FM-1
Section 2
NORMAL PROCEDURES
2
TABLE OF CONTENTS
Subject
Paragraph
Number
Page
Number
Introduction ............................................................................................
Cold Weather Operations.................................................................
Hot Weather Operations ...................................................................
Flight Planning .......................................................................................
Preflight Check.......................................................................................
Before Exterior Check ......................................................................
Exterior Check...................................................................................
Interior and Prestart Check...................................................................
Engine Start............................................................................................
Dry Motoring Run...............................................................................
Alternate Engine Start .......................................................................
Systems Check ......................................................................................
Preliminary Hydraulic Systems Check ............................................
FADEC Manual Check .......................................................................
Engine Runup.....................................................................................
Hydraulic Systems Check .................................................................
Before Takeoff........................................................................................
Takeoff ....................................................................................................
In-Flight Operations...............................................................................
Descent and Landing.............................................................................
Engine Shutdown...................................................................................
Postflight Check.....................................................................................
2-1 ...........
2-1-A .......
2-1-B .......
2-2 ...........
2-3 ...........
2-3-A .......
2-3-B .......
2-4 ...........
2-5 ...........
2-5-A .......
2-5-B .......
2-6 ...........
2-6-A .......
2-6-B .......
2-6-C .......
2-6-D .......
2-7 ...........
2-8 ...........
2-9 ...........
2-10 .........
2-11 .........
2-12 .........
2-3
2-3
2-3
2-3
2-3
2-4
2-4
2-7
2-9
2-10
2-10
2-11
2-11
2-12
2-12
2-12
2-13
2-13
2-14
2-14
2-15
2-16
Figure
Number
Page
Number
2-1 ...........
2-17
LIST OF FIGURES
Subject
Preflight Check Sequence.....................................................................
20 JUN 2007—Rev. 6———2-1/2-2
TEMPORARY REVISION
TC APPROVED
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
Section 2
BHT-407-FM-1
TR-10
NORMAL PROCEDURES
2
2-1. INTRODUCTION
Th is s ec tion c o ntain s in stru ct ion s an d
procedures for operating helicopter from
planning stage, through actual flight
conditions, to securing helicopter after
landing.
Normal and standard conditions are assumed
in these procedures. Pertinent data in other
sections is referenced when applicable.
Instructions and procedures contained herein
are written for purpose of standardization and
are not applicable to all situations.
2-1-A.
COLD WEATHER OPERATIONS
Battery starts have been demonstrated to
-29°C (-20°F) with standard 17 amp-hour
battery and -35°C (-31°F) with optional 28
amp-hour battery.
During engine start in cold temperatures,
initial engine oil pressure of 200 PSI and
pressure excursions down to 50 PSI during
warm up are normal. Normal oil pressure and
temperature indications as per Section 1
should be obtained after approximately 5
minutes at idle.
2-1-B.
HOT WEATHER OPERATIONS
CAUTION
DU R IN G EX T EN D ED H O VE R A T
TAKEOFF POWER WITH THE OAT
ABOVE 49.7°C (121.4°F), MONITOR
THE ENGINE OIL TEMPERATURE. IF
T E M P E R A T U R E
R I S E S ABNORMALLY, REDUCE
POWER OR TRANSITION TO
FORWARD
FLIGHT
UNTIL
TEMPERATURE DECREASES.
2-2. FLIGHT PLANNING
Each flight should be planned adequately to
ensure safe operations and to provide pilot
with data to be used during flight.
Check type of mission to be performed and
destination.
Determine that helicopter has adequate
performance to complete mission utilizing
appropriate performance charts in Section 4.
Determine that helicopter weight and balance
will be within limits during entire mission.
Utilize appropriate weight and balance charts
in Section 5 and limitations in Section 1.
2-3. PREFLIGHT CHECK
Pilot is responsible for determining whether
helicopter is in condition for a safe flight.
R e f e r t o F i g u re 2 -1 f o r p r e fl ig h t c h e c k
sequence.
NOTE
A preflight check is not intended to
be a detailed mechanical inspection,
but simply a guide to help pilot check
condition of helicopter. It may be as
comprehensive as conditions
warrant at discretion of pilot.
All areas checked shall include a
visual check for evidence of
corrosion, particularly when
helicopter is flown near salt water or
in areas of high industrial emissions.
19 FEB 2007—Rev. 5———2-3
BHT-407-FM-1
2-3-A.
TC APPROVED
BEFORE EXTERIOR CHECK
4.
Transmission — Check oil level.
Verify actual presence of oil in sight
gauge.
5.
Transmission oil cooler lines —
Condition and security.
6.
Transmission mounts — Condition
and security.
1.
Flight planning — Completed.
2.
Publications — Checked.
3.
GW and CG — Computed.
4.
Helicopter servicing — Completed.
5.
Battery — Connected.
7.
Main driveshaft — Condition.
EXTERIOR CHECK
8.
Access door — Secured.
9.
Fuel filler cap — Visually check fuel
level and cap secured.
2-3-B.
2-3-B-1.
FUSELAGE — CABIN RIGHT SIDE
NOTE
WARNING
FAILURE TO REMOVE ROTOR
TIEDOWNS BEFORE ENGINE
STARTING MAY RESULT IN SEVERE
DAMAGE AND POSSIBLE INJURY.
1.
All main rotor blades — Tiedowns
removed, condition.
2.
Right static port — Condition.
3.
If helicopter is not parked on a level
surface, fuel sump may not properly
drain contaminants.
10. Fuel sump — Drain fuel sample as
follows:
a.
RIGHT and LEFT FUEL BOOST/
XFR circuit breaker switches —
OFF.
Cabin doors and hinge bolts —
Condition and security.
b.
BATT switch — BATT (on).
c.
FUEL VALVE switch — OFF.
4.
Windows — Condition and security.
d.
5.
Landing gear — Condition. Ground
handling wheel removed.
FWD and AFT FUEL SUMP drain
buttons — Press, drain sample,
then release.
6.
Forward and aft crosstube fairings (if
installed) — Secured, condition, and
aligned.
2-3-B-2.
FUSELAGE — CENTER RIGHT
SIDE
1.
Engine inlet — Condition; remove
inlet covers.
2.
Cabin roof, transmission cowling,
and engine air inlet area — Cleaned
of all debris, accumulated snow and
ice; cowling secured.
3.
Forward fairing — Secured.
2-4———Rev. 5—19 FEB 2007
11. Airframe fuel filter — Drain and
check before first flight of day as
follows:
a.
RIGHT and LEFT FUEL BOOST/
XFR circuit breaker switches —
LEFT and RIGHT (on).
b.
FUEL VALVE switch — ON.
c.
Fuel filter drain valve — Open,
drain sample, then close.
12. Fuel filter test switch — Press and
check FUEL FILTER caution light
illuminates. Release switch and
check light extinguishes.
TR-10
TEMPORARY REVISION
TC APPROVED
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
2-3-B-4.
13. FUEL VALVE switch — OFF.
14. LEFT and RIGHT FUEL BOOST/XFR
circuit breaker switches — OFF.
15. BATT switch — OFF.
16. Power plant area:
a.
Main driveshaft aft flexure —
Condition.
b.
Engine — Condition, security of
attachments, evidence of oil
leakage.
c.
Engine mounts — Condition and
security.
d.
Throttle linkage — Condition,
security,
and
freedom
of
operation.
e.
BHT-407-FM-1
1.
Vertical fin — Condition.
2.
Tail rotor guard — Condition and
security.
3.
Anticollision light — Condition and
security of lens.
4.
Aft position light — Condition.
5.
Tail rotor gearbox — Oil level, leaks
and security.
6.
Tail rotor — Tiedown removed,
condition and free movement.
7.
Tail rotor controls — Condition and
security.
8.
Tail rotor blades:
Engine fuel pump — Security
and condition, evidence of
leakage.
f.
H y d r o m e c h a n i c a l u n it —
S e c u r i ty a n d c o n d i ti o n ,
evidence of leakage.
g.
Hoses and tubing — Chafing,
security, and condition.
9.
17. Engine cowl — Secured.
18. Generator cooling scoop — Clear of
debris.
19. Oil tank — Leaks, security, and cap
secured.
2-3-B-5.
2-3-B-3.
FUSELAGE — AFT RIGHT SIDE
1.
Fuselage — Condition.
2.
Tail rotor driveshaft
Condition and security.
3.
Tailboom — Condition.
4.
Horizontal stabilizer and position
light — Condition and security.
cover
—
a.
General condition.
b.
Tip block — Security and seal
integrity.
c.
Internal blade root — Clear of
snow and ice.
Tail rotor yoke — Condition,
evidence of static stop contact
damage (deformed static stop yield
indicator).
FUSELAGE — AFT LEFT SIDE
1.
Tailboom — Condition.
2.
Tail rotor driveshaft
Condition and security.
3.
Horizontal stabilizer area:
20. Access door — Secured.
21. Aft fairing — Secured.
FUSELAGE — FULL AFT
cover
—
a.
Horizontal stabilizer — General
condition and security of
attachment.
b.
Position light — Condition and
security.
c.
Forward and aft section of left
upper stabilizer support to
tailboom area — Condition of
tailboom.
19 FEB 2007—Rev. 5———2-5
TR-10
TEMPORARY REVISION
BHT-407-FM-1
4.
Fuselage — Condition.
5.
Forward
tail
rotor
driveshaft
coupling — Condition of splined
adapter.
6.
Oil cooler blower shaft hanger
bearings — Evidence of grease
leakage and overheating.
7.
TC APPROVED
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
Oil cooler blower — Clear
obstructions and condition.
j.
Rotor brake disc and caliper (if
installed) — Condition, security
of attachment and leakage.
Ensure brake pads are retracted
from brake disc.
k.
Engine cowling — Secured.
l.
Air
induction
Secured.
of
8.
Oil cooler — Condition and leaks.
9.
Oil cooler blower access door —
Secured.
10. Oil tank sight glass — Check oil level.
cowling
—
m. Cabin
roof,
transmission
cowling, engine air inlet area,
and plenum — Clear of all
debris, accumulated snow and
ice; cowling secured.
15. Transmission area:
11. Aft fairing — Secured.
12. Baggage compartment — Cargo tied
down, door secured.
a.
Transmission
mounts
Condition and security
elastomeric mounts.
b.
Transmission oil filter — Ensure
bypass indicator not extended.
c.
Main driveshaft — Condition.
13. Exhaust cover — Removed.
14. Power plant area:
—
of
a.
Engine — Condition, security of
attachments.
d.
Transducers and pressure lines
— Condition and security.
b.
Engine mounts — Condition and
security.
e.
Access door — Secured.
c.
Exhaust stack — Condition and
security.
d.
Evidence of fuel and oil leaks.
e.
Fuel and oil filter bypass
indicators — Check retracted.
f.
Hoses and tubing for chafing
and condition.
g.
Pneumatic lines — Condition
and security.
h.
Tail
rotor
driveshaft
Condition of
splines
couplings.
i.
Air induction diffuser duct —
Condition and security.
2-6———Rev. 5—19 FEB 2007
2-3-B-6.
CABIN ROOF
1.
Main rotor dampers and fairing —
Condition and security.
2.
Main rotor hub, yoke and frahm —
Condition and security.
3.
Main rotor
Condition.
4.
Pitch horn bearing — Wear and
security.
5.
Main rotor pitch links — Condition
and security of attachment bolts and
locking hardware.
6.
Swashplate assembly — Condition,
security of attached controls, and
boot condition.
—
and
blade
and
skin
—
TC APPROVED
BHT-407-FM-1
7.
Control linkages to swashplate —
Condition, security of attachment
bolts and locking hardware.
8.
Control tube hydraulics-off balance
springs — Condition and security.
9.
Hydraulic reservoir filler cap —
Closed and locked.
7.
External power door — Condition
and security.
8.
Landing light lamps — Condition.
9.
Antennas — Condition and security.
2-4. INTERIOR
CHECK
AND
PRESTART
10. Hydraulic system filters — Bypass
indicator retracted.
1.
Cabin interior — Clean, equipment
secured.
11. Hydraulic actuators and lines —
Condition, security, interference,
leakage.
2.
Fire extinguisher — Installed and
secured.
3.
Cabin loading — Maintain CG within
limits.
4.
Passenger seat belts — Secured.
5.
Copilot seat belt — Secured (if solo).
6.
Doors — Secured.
7.
Throttle — Closed.
8.
LDG LTS switch — OFF.
9.
Communications switches — Set.
2-3-B-7.
1.
FUSELAGE — CABIN LEFT SIDE
Forward fairing and access door —
Secured.
2.
Cabin doors and hinge bolts —
Condition and security.
3.
Windows — Condition and security.
4.
Hydraulic reservoir — Check fluid
level.
5.
Landing gear — Condition and
ground handling wheel removed.
10. Altimeter — Set.
6.
Forward and aft crosstube fairings (if
installed) — Secured, condition, and
aligned.
12. Overhead switches — Set:
7.
2-3-B-8.
Left static port — Condition.
FUSELAGE — FRONT
1.
Exterior surfaces — Condition.
2.
Windshield
cleanliness.
3.
Battery and vent lines — Condition
and security.
4.
HOUR METER circuit breaker — In.
5.
Battery access door — Secured.
6.
Pitot tube — Cover removed, clear of
obstructions.
—
Condition
and
11. Instruments — Correct indications.
a.
BATT switch — OFF.
b.
GEN switch — OFF.
c.
PART SEP switch (if installed) —
OFF.
d.
ANTI COLL LT switch — ANTI
COLL LT (on).
e.
HYD SYS switch — HYD SYS
(on).
f.
CABIN LT/PASS switch — OFF.
g.
POS LT switch — As desired.
h.
DEFOG switch — OFF.
i.
PITOT HEATER switch — OFF.
19 FEB 2007—Rev. 5———2-7
BHT-407-FM-1
TC APPROVED
j.
ENG ANTI ICE switch — OFF.
k.
AVIONICS MASTER switch —
OFF.
l.
HEATER switch (if installed) —
OFF.
m. INSTR LT rheostat — OFF.
13. Overhead circuit breaker switches —
OFF.
14. Overhead circuit breakers — In.
15. Rotor brake handle (if installed) — Up
and latched.
CAUTION
c.
After 3 seconds; ENG OUT,
FADEC
DEGRADE,
FADEC
FAULT, RESTART FAULT, and
ENGINE
OVSPD
lights
illuminate with activation of
engine out audio for 3 seconds.
d.
ENG OUT light re-illuminates
with reactivation of engine out
audio, after 3 seconds.
18. HORN MUTE button — Press to mute.
19. Caution lights — ENG OUT, XMSN
OIL PRESS, RPM, HYDRAULIC
SYSTEM, GEN FAIL, L/FUEL BOOST,
R/FUEL BOOST, L/FUEL XFR, and R/
FUEL XFR will be illuminated.
NOTE
28 VDC GPU SHALL BE 500
AMPERES OR LESS TO REDUCE
RISK OF STARTER DAMAGE FROM
OVERHEATING.
L/FUEL XFR and R/FUEL XFR will not
be illuminated when forward fuel
tank is empty.
16. GPU — Connected (if used).
20. PEDAL
STOP
annunciator:
17. BATT switch — ON for battery start,
ON for GPU start, OFF for battery cart
start. Observe the following:
a.
Low rotor audio horn activated.
NOTE
With “Ahlers” NR/NP gauge
installed, NR/NP needles do not self
test.
b.
For 8 seconds,
(1) Trend
arcs
on
LCD
instruments indicate full
scale.
(2) TORQUE and NG digits
display 8188.8.
(3) MGT and FUEL
display 81888.
digits
(4) NR and NP needles move to
107%
and
100%,
respectively.
2-8———Rev. 7—30 JUL 2008
PTT
switch
Pedals — Centered.
Press — Verify PEDAL STOP caution
and
ENGAGED
annunciator
illuminated and left pedal travel
restricted.
Release — Verify PEDAL STOP
caution and ENGAGED annunciator
extinguished and both pedals travel
unrestricted.
21. Flight controls — Loosen frictions;
check travel and verify CYCLIC
CENTERING light operation; position
for start. Tighten friction as desired.
22. Throttle — Check freedom of travel
and appropriate operation at OFF, I
(idle), FLY and MAX positions.
Return throttle to OFF position.
NOTE
With INSTR LT rheostat on and CAUT
LT switch positioned to DIM, caution
TR-11
TEMPORARY REVISION
THIS PAGE HAS
AN ASSOCIATED
TEMPORARY REVISION.
TC APPROVED
lights are dimmed to a fixed intensity
and cannot be adjusted by INSTR LT
rheostat.
BHT-407-FM-1
2.
23. INSTR LT rheostat — As desired.
24. CAUT LT switch — As desired.
25. FUEL BOOST/XFR circuit breaker
switches — LEFT (on) and RIGHT
(on) and verify all boost and transfer
caution lights extinguish.
Cyclic and pedals — Centered and
CYCLIC
CENTERING
light
extinguished.
NOTE
If throttle is positioned in idle for
more than 60 seconds, starter
latching is disabled and throttle must
be repositioned to cut off and then
back to idle to enable it for another 60
seconds.
26. FUEL pressure — Check.
30. FADEC HORN TEST button — Press
to test.
It is recommended that MGT be
below 150°C when below 10,000 feet
HP or below 65 °C when above 10,000
feet HP prior to attempting an engine
start. Compliance with this
recommendation will allow for cooler
starts and reduce potential of
reaching hot start abort limits. Refer
to DRY MOTORING RUN, paragraph
2-5-A.
31. FADEC MODE switch — AUTO.
3.
Throttle — Idle position.
4.
START switch — Momentarily press
(hold for approximately 1 second)
and observe START and AUTO
RELIGHT lights are illuminated.
5.
MGT — Monitor.
27. CAUTION LT TEST button — Press to
test.
28. INSTR CHK button — Press and
check for exceedances.
29. LCD TEST button — Press to test, if
desired.
32. FUEL VALVE switch — ON, guard
closed,
FUEL
VALVE
light
illuminates then extinguishes.
33. FUEL QTY — Check TOTAL and FWD
tank quantity.
34. OAT/VOLTS display — Check OAT
and select VOLTS.
CAUTION
CAUTION
ANY ATTEMPT TO START ENGINE
WHEN VOLTAGE IS BELOW 24
V O L T S M A Y R E S U L T IN A H O T
START. MONITOR FOR FADEC
FAILURE. IF FADEC FAILS (FADEC
FAIL WARNING LIGHT), ABORT
START BY ROLLING THROTTLE TO
CUTOFF AND ENGAGE STARTER
TO REDUCE MGT.
2-5. ENGINE START
1.
Collective — Full down.
IF MAIN ROTOR IS NOT ROTATING
B Y 2 5% N G , A BO R T ST A R T B Y
ROLLING THROTTLE TO CUTOFF.
ENSURE
STARTER
HAS
DISENGAGED
WHEN
MGT
DECREASES BELOW 150°C.
6.
START light — Extinguished at 50%
NG (starter has disengaged).
7.
AUTO RELIGHT light — Extinguished
at 60% NG.
8.
ENG and XMSN OIL pressures —
Check.
30 JUL 2008—Rev. 7———2-9
BHT-407-FM-1
TC APPROVED
2-5-A.
DRY MOTORING RUN
CAUTION
IF ENGINE HAS BEEN SHUT DOWN
F O R M O R E TH A N 15 M IN U T E S,
STABILIZE AT IDLE FOR 1 MINUTE
BEFORE INCREASING THROTTLE.
The following procedure is used to reduce
residual MGT to recommended levels for
engine start.
1.
Throttle — Closed position.
2.
START switch — Hold engaged for 15
seconds, then release.
NOTE
During cold temperature operations,
normal transmission and engine oil
pressure limits may be exceeded
during start. Stabilize engine at idle
until minimum temperature and
pressure limits are attained.
9.
Idle — 63 ±1% NG.
10. BATT switch — ON (if applicable).
11. GPU — Disconnect and close door (if
applicable).
Follow ENGINE START procedure, paragraph
2-5, once 0% NG is indicated.
2-5-B.
ALTERNATE ENGINE START
This procedure may be used in hot and/or
high altitude environment where aborted hot
starts have been experienced and when prior
troubleshooting has not revealed any engine
maintenance issues.
1.
Collective — Full down.
2.
Cyclic and pedals — Centered and
CYCLIC
CENTERING
light
extinguished.
12. GEN switch — GEN (on); observe
GEN FAIL light extinguishes.
NOTE
NOTE
Turn gen erator O FF if ammeter
indication drops to zero amps after
an initial full scale indication. One
reset is allowed. RESET generator
and then turn generator back ON
(applicable with AMPS/FUEL PSI
gauge PN 407-075-024-101 and sub.).
Refer to BHT-407-MD-1.
13. Voltmeter — 28.5 ±0.5 volts.
14. FLIGHT INSTR circuit breaker
switches (3) (if installed) — DG, ATT
and TURN (on).
It is recommended that MGT be
below 150°C when below 10,000 feet
HP or below 65°C when above 10,000
feet HP prior to attempting an engine
start. Compliance with this
recommendation will allow for cooler
starts and reduce potential of
reaching hot start abort limits. Refer
to DRY MOTORING RUN, paragraph
2-5-A.
3.
Throttle — Closed position.
4.
START switch — Hold engaged and
observe START and AUTO RELIGHT
lights are illuminated.
NOTE
If dual controls are installed, guard
t h r o t tl e to p r e v e n t i n a d v e r t e n t
manipulation from co-pilot position.
2-10———Rev. 5—19 FEB 2007
a.
5.
Throttle — Open to IDLE at
approximately 16% NG.
MGT — Monitor.
TC APPROVED
BHT-407-FM-1
NOTE
Engine will detect light off and
smoothly accelerate to idle while
limiting MGT, if necessary.
a.
START switch — Once light off
is detected, release.
CAUTION
IF MAIN ROTOR IS NOT ROTATING
B Y 2 5% N G , A BO R T S T A R T B Y
ROLLING THROTTLE TO CUTOFF.
ENSURE
STARTER
HAS
DISENGAGED
WHEN
MGT
DECREASES BELOW 150°C.
6.
START light — Extinguished at 50%
NG (starter has disengaged).
7.
AUTO RELIGHT light — Extinguished
at 60% NG.
8.
ENG and XMSN OIL pressures —
Check.
11. GPU — Disconnect and close door (if
applicable).
12. GEN switch — GEN (on); observe
GEN FAIL light extinguishes.
NOTE
Turn generator OFF if ammeter
indication drops to zero amps after
an initial full scale indication. One
reset is allowed. RESET generator
and then turn generator back ON
(applicable with AMPS/FUEL PSI
g a u g e P N 4 0 7 -0 7 5 - 02 4 - 1 01 a n d
subsequent). Refer to BHT-407-MD-1.
13. Voltmeter — 28.5 ±0.5 volts.
14. FLIGHT INSTR circuit breaker
switches (3) (if installed) — DG, ATT
and TURN (on).
NOTE
If dual controls are installed, guard
t h r o t t le to p r e v e n t in a d v e r t e n t
manipulation from co-pilot position.
2-6. SYSTEMS CHECK
CAUTION
IF ENGINE HAS BEEN SHUT DOWN
F O R M O R E TH A N 15 M IN U T E S,
STABILIZE AT IDLE FOR 1 MINUTE
BEFORE INCREASING THROTTLE.
NOTE
During cold temperature operations,
normal transmission and engine oil
pressure limits may be exceeded
during start. Stabilize engine at idle
until minimum temperature and
pressure limits are attained.
9.
Idle — 63 ±1% NG.
10. BATT switch — ON (if applicable).
2-6-A.
PRELIMINARY HYDRAULIC
SYSTEMS CHECK
NOTE
Uncommanded control movement or
motoring with hydraulic system off
m a y in d i c a te h y d r a u lic s y s t e m
malfunction.
1.
HYD SYS switch — OFF.
2.
HYDRAULIC SYSTEM caution light
— Illuminated.
3.
HYD SYS switch — HYD SYS (on).
4.
HYDRAULIC SYSTEM caution light
— Extinguished.
19 FEB 2007—Rev. 5———2-11
BHT-407-FM-1
2-6-B.
TC APPROVED
FADEC MANUAL CHECK
2.
WARNING
AUTO TO MANUAL MODE
TRANSITIONS WITH NR/NP AT 100%
FLAT PITCH CAN RESULT IN RAPID
NR/NP ACCELERATION IN
APPROXIMATELY 7 SECONDS. TO
A V O ID P O S S I B L E O V E R S P E E D
CONDITION, PERFORM THE
FOLLOWING CHECK AT IDLE (63%
NG).
1.
Throttle — Idle (63% NG).
2.
FADEC MODE switch — MAN.
3.
FADEC
MANUAL
and
AUTO
RELIGHT lights — Illuminated.
4.
Check NG stabilized at 75% or less.
5.
Throttle — Increase slowly to ensure
engine responds, then return to idle.
6.
FADEC MODE switch - AUTO.
7.
FADEC
MANUAL
and
AUTO
RELIGHT lights — Extinguished.
2-6-C.
ENGINE RUNUP
NOTE
Overhead
circuit
breakers
highlighted with arrow graphic
;
are powered through AVIONICS
MASTER switch.
3.
AVIONICS MASTER switch
AVIONICS MASTER (on).
4.
ELT (if installed) — Check for
inadvertent transmission.
5.
Flight controls — Check freedom
with minimum friction.
6.
ENG ANTI ICE switch — ENG ANTI
ICE (on); check for MGT increase and
illumination of ENGINE ANTI-ICE
light (if installed).
7.
ENG ANTI ICE switch — OFF; check
MGT returns to normal and ENGINE
ANTI-ICE
light
(if
installed)
extinguishes; then ENG ANTI ICE
(on) if required.
If temperature is below 5°C (40°F)
and visible moisture is present, ENG
ANTI ICE shall be on.
CAUTION
1.
Throttle — Increase smoothly to FLY
detent position while maintaining
torque below 40%. Check RPM
warning light extinguished at 95%
NR.
2-12———Rev. 6—20 JUN 2007
—
NOTE
8.
FAILURE
TO
SMOOTHLY
TRANSITION THE THROTTLE FROM
THE GROUND IDLE POSITION TO
THE FLY POSITION, MAINTAINING
TO R Q U E L E S S T H A N 4 0% M A Y
RESULT IN ENGINE OVERSPEED OR
OVERTORQUE.
NR and NP needles — Check
matching and indicating 100%.
2-6-D.
PART SEP switch (if installed) — As
required.
HYDRAULIC SYSTEMS CHECK
NOTE
Hydraulic systems check is to
determine proper operation of
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.
1.
Collective — Full down.
TC APPROVED
BHT-407-FM-1
2.
NR — 100% RPM.
3.
HYD SYS switch — OFF.
4.
HYDRAULIC SYSTEM caution light
— Illuminated.
5.
Cyclic — Centered.
6.
Cyclic control — Check normal
operation by moving cyclic forward
and aft, then left and right
(approximately 1 inch). Center cyclic.
CAUTION
FAILURE TO POSITION AND
MAINTAIN THROTTLE IN FLY
DETENT POSITION PRIOR TO
TAKEOFF AND D U R IN G N O R M A L
F L I G H T OPERATIONS CAN LIMIT
AVAILABLE ENGINE POWER.
6.
Throttle — Open to FLY detent
position. Check 99 to 100% NR/NP.
Collective — Check normal
operation by increasing collective
slightly (1 to 2 inches). Repeat two to
three times as required. Return to full
down position.
7.
Engine, transmission, and electrical
instruments — Within limits.
8.
Flight and navigation instruments —
Check.
8.
Pedals — Check normal operation by
displacing pedals slightly (1 inch).
9.
FUEL QTY — Note indication.
9.
HYD SYS switch — HYD SYS (on).
10. FUEL QTY FWD TANK button —
Press, note fuel remaining in forward
cell.
7.
10. HYDRAULIC SYSTEM caution light
— Extinguished.
11. Cyclic and collective friction — Set
as desired.
2-8. TAKEOFF
1.
Rear facing seat headrests
Adjusted to proper position.
—
2-7. BEFORE TAKEOFF
NOTE
1.
ENG ANTI ICE switch — As required.
2.
Light switches — As required.
3.
INSTR LT rheostat — As desired.
NOTE
For night flight, it is recommended to
point the map light at the flight
instruments and set to a low
intensity. Sufficient night lighting will
b e p r o v id e d i n t h e e v e n t o f a n
instrument lighting failure.
4.
Radio(s) — Check as required.
5.
Flight controls — Position and adjust
frictions for takeoff.
During takeoffs disregard CYCLIC
CENTERING light and position cyclic
as required.
2.
Collective — Increase to hover.
3.
Directional control — As required to
maintain desired heading.
4.
Cyclic — Apply as required to
accelerate smoothly.
5.
Increase collective, up to 5% torque
above hover power, to obtain desired
rate of climb and airspeed. Once
clear of the HV diagram shaded
areas, adjust power and airspeed as
desired.
20 JUN 2007—Rev. 6———2-13
BHT-407-FM-1
6.
TC APPROVED
PEDAL STOP PTT switch — Check
ENGAGED annunciator illuminated
above 55 ±5 KIAS.
6.
2-9. IN-FLIGHT OPERATIONS
1.
AIRSPEED — As desired (not to
exceed VNE at flight altitude).
CAUTION
AT HIGH POWER AND HIGH
AIRSPEED, CYCLIC ONLY
ACCELERATIONS AND
MANEUVERING
MAY
SIGNIFICANTLY INCREASE MGT
AND
TORQUE
WITH
NO
COLLECTIVE
INPUT.
THIS
INCREASE IS MORE RAPID AT
LOWER OAT.
NOTE
Pilot shall keep feet on tail rotor
pedals at all times. Do not press
PEDAL STOP PTT switch in flight.
2.
3.
4.
PEDAL STOP PTT switch — Check
ENGAGED annunciator illuminated
above 55 ±5 KIAS.
ENG ANTI ICE and PITOT HEATER
switches — ENG ANTI ICE and PITOT
HEATER switches on in visible
moisture when ambient temperature
is at or below 5°C (40°F).
PITOT HEATER — Confirm operation
(increase ammeter load).
NOTE
When ENG ANTI ICE switch is in ENG
ANTI ICE (on), MGT will increase.
Monitor MGT when selecting ENG
ANTI ICE at high power settings.
5.
Altimeter — Within limits.
2-14———Rev. 6—20 JUN 2007
FUEL QTY FWD TANK button —
Press, note forward fuel tank
indication.
NOTE
F u ll fo r w a rd fu e l ta n k qu a n ti ty
(approximately 256 pounds) will be
ind ica ted a t ap pro xima tely 7 70
pounds or greater total fuel. Fuel
transfer will be complete at
approximately 195 pounds total fuel.
2-10. DESCENT AND LANDING
NOTE
Large reductions in collective pitch
at hea vy GW may permit NR to
increase independent of NP (needles
split). Main rotor may be reengaged
with a smooth increase in collective
pitch.
1.
Rear facing seat headrests
Adjusted to proper position.
—
2.
Flight controls — Adjust friction as
desired.
3.
Throttle — Fly detent position. Check
99 to 100% NP.
4.
Flight path — As required for type of
approach.
5.
ENG ANTI ICE — As required.
6.
LDG LTS switch — As desired.
NOTE
During run-on or slope landings,
disregard CYCLIC CENTERING light
and position cyclic as required. After
landing is completed and collective
is full down, reposition cyclic so that
CYCLIC CENTERING light is
extinguished.
7.
PEDAL STOP PTT switch — Check
ENGAGED annunciator extinguished
below 50 ±5 KIAS.
TC APPROVED
BHT-407-FM-1
2-11. ENGINE SHUTDOWN
12. AVIONICS MASTER switch — OFF.
13. GEN switch — OFF.
1.
Collective — Full down.
2.
Cyclic and pedals — Centered and
CYCLIC
CENTERING
light
extinguished.
3.
Cyclic friction — Increase so that
cyclic maintains centered position.
4.
LDG LTS switch — OFF.
5.
Throttle — Reduce to idle stop.
Check RPM warning light illuminated
and audio on at 95% NR.
NOTE
Overspeed shutdown test should be
accomplished on first engine
shutdown of the day. ENGINE
OVSPD light will momentarily
illuminate in addition to those lights
that illuminate during a normal
shutdown.
15. IDLE REL switch — Press and hold.
NOTE
If dual controls are installed, guard
t h r o t tl e to p r e v e n t i n a d v e r t e n t
manipulation from co-pilot position.
6.
HORN MUTE button — Press to mute.
7.
MGT — Stabilize at idle for 2 minutes.
8.
ENG ANTI ICE switch — OFF.
9.
FLIGHT INSTR circuit breakers
switches (if installed) — OFF.
10. FUEL BOOST/XFR LEFT
breaker switch — OFF.
14. OVSPD TEST button — If required;
press, hold 1 second, and release.
circuit
NOTE
Left fuel boost and transfer pumps
will continue to operate until either
LE F T F U E L B O O S T/ XF R cir cu it
breaker switch (highlighted with
ye ll ow b or de r) o r F UE L VA L V E
switch is positioned to OFF. These
pumps operate directly from battery
and will not be deactivated when
BATT switch is OFF. Battery power
will be depleted if both switches
remain on.
11. ELT (if installed) — Check for
inadvertent transmission.
CAUTION
POSITIONING THROTTLE OUT OF
CUT-OFF DURING NG SPOOL DOWN
MAY CAUSE POST ENGINE
SHUTDOWN FIRE.
16. Throttle — Closed; check MGT and
NG
decreasing, ENGINE
OUT
warning light illuminated and audio
on at 55 ±1%.
17. HORN MUTE button — Press to mute.
CAUTION
AVOID RAPID ENGAGEMENT OF
ROTOR BRAKE IF HELICOPTER IS
ON ICE OR OTHER SLIPPERY OR
LOOSE SURFACE TO PREVENT
ROTATION OF HELICOPTER.
18. Rotor brake (if installed) — Apply full
rotor brake at or below 40% NR.
Return rotor brake handle to stowed
position just prior to main rotor
stopping.
19. FUEL VALVE switch — OFF.
20 JUN 2007—Rev. 6———2-15
BHT-407-FM-1
TC APPROVED
CAUTION
DO NOT INCREASE COLLECTIVE OR
APPLY LEFT TAIL ROTOR PEDAL
TO SLOW ROTOR DURING
COASTDOWN.
20. Pilot — Remain on flight controls
until rotor has come to a complete
stop.
21. All overhead switches, except HYD
SYS switch — OFF.
NOTE
pounds total fuel quantity, up to 18
pounds of fuel may remain in forward
fuel cell as unusable.
2-12. POSTFLIGHT CHECK
If any of following conditions exist:
•
Thunderstorms are in local area or
forecasted.
•
Winds in excess of 35 knots or a
gust spread of 15 knots exists or is
forecasted.
•
Helicopter is parked within 150 feet
of hovering or taxiing aircraft that
are in excess of basic GW of
helicopter.
•
Helicopter to be left unattended.
Ensure engine rotation has
completely stopped prior to
positioning BATT switch to OFF.
22. BATT switch — OFF, with NG at 0%.
Perform following:
CAUTION
APPLICABLE MAINTENANCE
ACTION MUST BE PER FO RMED
PRIOR TO FURTHER FLIGHT IF A
FADEC LIGHT HAS ILLUMINATED
DURING THE PREVIOUS FLIGHT OR
ON ENGINE SHUTDOWN.
NOTE
If shutting down at, or refueling to,
between approximately 195 to 213
2-16———Rev. 7—30 JUL 2008
1.
Install main rotor blade tie-downs.
2.
Secure tail rotor loosely to tailboom
with tie-down strap to prevent
excessive flapping.
3.
Install exhaust cover, engine inlet
protective plugs and pitot cover.
NOTE
Refer to BHT-407-MD-1 for additional
tie-down data.
TC APPROVED
BHT-407-FM-1
Figure 2-1. Preflight check sequence
17 DEC 2002
2-17/2-18
TC APPROVED
BHT-407-FM-1
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3
TABLE OF CONTENTS
Subject
Paragraph
Number
Page
Number
Introduction ............................................................................................
Definitions ..............................................................................................
Engine .....................................................................................................
Engine Failure ....................................................................................
Engine Restart in Flight.....................................................................
Engine Underspeed ...........................................................................
Engine Overspeed .............................................................................
Engine Compressor Stall ..................................................................
Engine Hot Start/Shutdown ..............................................................
Engine Oil Pressure Low or Fluctuating..........................................
Engine Oil Temperature High ...........................................................
Driveshaft Failure...............................................................................
FADEC Failure....................................................................................
Fire ..........................................................................................................
Engine Fire on Ground ......................................................................
Engine Fire During Flight ..................................................................
Cabin Smoke or Fumes .....................................................................
Tail Rotor ................................................................................................
Complete Loss of Tail Rotor Thrust .................................................
Fixed Pitch Failures ...........................................................................
Hydraulic System...................................................................................
Loss of Hydraulic Pressure ..............................................................
Flight Control Actuator Malfunction ................................................
Electrical System ...................................................................................
Generator Failure ...............................................................................
Excessive Electrical Load .................................................................
Fuel System............................................................................................
Cyclic Jam ..............................................................................................
Warning, Caution, and Advisory Lights/Messages ............................
3-1 ...........
3-2 ...........
3-3 ...........
3-3-A .......
3-3-B .......
3-3-C .......
3-3-D .......
3-3-E .......
3-3-F........
3-3-G .......
3-3-H .......
3-3-J........
3-3-K .......
3-4 ...........
3-4-A .......
3-4-B .......
3-4-C .......
3-5 ...........
3-5-A .......
3-5-B .......
3-6 ...........
3-6-A .......
3-6-B .......
3-7 ...........
3-7-A .......
3-7-B .......
3-8 ...........
3-9 ...........
3-10 .........
3-3
3-3
3-3
3-3
3-4
3-6
3-6
3-6
3-7
3-7
3-7
3-8
3-8
3-9
3-9
3-9
3-9
3-10
3-10
3-10
3-11
3-11
3-12
3-12
3-12
3-12
3-13
3-13
3-14
Table
Number
Page
Number
3-1 ...........
3-2 ...........
3-15
3-16
LIST OF TABLES
Subject
Warning (Red) Lights.............................................................................
Caution (Amber) and Advisory (White/Green) Lights.........................
30 JUL 2008
Rev. 7
3-1/3-2
TC APPROVED
BHT-407-FM-1
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3
3-1.
INTRODUCTION
Following procedures contain indications of
failures or malfunctions which affect safety of
c r e w, h e l i c o p t e r, g r o u n d p e r s o n n e l o r
property; use of emergency features of
primary and backup systems; and appropriate
warnings, cautions, and explanatory notes.
Tables 3-1 and 3-2 list fault conditions and
corrective actions for warning lights and
caution/advisory lights respectively.
LAND AS SOON
AS POSSIBLE
Land without delay at
nearest suitable area
(i.e., open field) at which
a safe approach and
landing is reasonably
assured.
LAND AS SOON
AS PRACTICAL
Landing site and
duration of flight are at
discretion of pilot.
Extended flight beyond
nearest approved
landing area is not
recommended.
NOTE
All corrective action procedures
listed herein assume pilot gives first
priority to helicopter control and a
safe flight path.
A tripped circuit breaker should not
be reset in flight unless deemed
necessary for safe completion of the
flight.
If a tripped circuit breaker is deemed
necessary for safe completion of the
flight, it should only be reset one
time.
Helicopter should not be operated following
any precautionary landing until cause of
malfunction has been determined and
corrective maintenance action taken.
3-2.
DEFINITIONS
Following terms indicate degree of urgency in
landing helicopter.
Following terms are used to describe
operating condition of a system, subsystem,
assembly, or component.
Affected
Fails to operate in
intended or usual
manner.
Normal
Operates in intended
or usual manner.
3-3.
3-3-A.
3-3-A-1.
ENGINE
ENGINE FAILURE
ENGINE FAILURE — HOVERING
INDICATIONS:
1.
Left yaw.
2.
ENGINE OUT and RPM warning
lights illuminated.
30 JUL 2008
Rev. 7
3-3
BHT-407-FM-1
TC APPROVED
3.
Engine instruments indicate power
loss.
4.
Engine out audio activated when NG
drops below 55%.
5.
2.
Maintain
control.
heading
and
attitude
Maintain
control.
2.
Collective — Adjust as required to
maintain 85 to 107% NR.
3.
Land.
4.
Shut down helicopter.
ENGINE FAILURE — INFLIGHT
Left yaw.
2.
ENGINE OUT and RPM warning
lights illuminated.
3.
Engine instruments indicate power
loss.
4.
Engine out audio activated when NG
drops below 55%.
5.
NR decreasing with RPM warning
light and audio on when NR drops
below 95%.
attitude
Cyclic — Adjust to obtain desired
autorotative AIRSPEED.
NOTE
Maximum AIRSPEED for steady state
autorotation is 100 KIAS. Minimum rate
of descent airspeed is 55 KIAS.
Maximum glide distance airspeed is 80
KIAS.
4.
Attempt engine restart if ample
altitude remains. (Refer to ENGINE
RESTART, paragraph 3-3-B).
5.
FUEL VALVE switch — OFF.
6.
At low altitude:
INDICATIONS:
1.
and
Maintaining NR at high end of
operating range will provide maximum
rotor energy to accomplish landing, but
will cause an increased rate of descent.
Collective — Adjust to control NR
and rate of descent. Increase prior to
ground contact to cushion landing.
Amplitude of collective movement is a
function of height above ground. Any
forward airspeed will aid in ability to
cushion landing.
heading
NOTE
3.
NOTE
3-3-A-2.
1.
NR decreasing with RPM warning
light and audio on when NR drops
below 95%.
PROCEDURE:
1.
PROCEDURE:
a.
Throttle — Closed.
b.
Flare to lose airspeed.
7.
Apply collective as flare effect
decreases to further reduce forward
speed and cushion landing. Upon
ground contact, collective shall be
reduced smoothly while maintaining
cyclic in neutral or centered
position.
8.
Complete helicopter shutdown.
3-3-B.
ENGINE RESTART IN FLIGHT
An engine restart may be attempted in flight if
time and altitude permit.
3-4
17 DEC 2002
TC APPROVED
BHT-407-FM-1
3-3-B-2.
RESTART — MANUAL MODE
CAUTION
IF CAUSE OF FAILURE IS OBVIOUSLY
MECHANICAL, AS EVIDENCED BY
ABNORMAL METALLIC OR GRINDING
SOUNDS, DO NOT ATTEMPT A
RESTART.
3-3-B-1.
R E S TA R T FA U LT O R FA D E C M A N U A L
LIGHTS ILLUMINATED.
l PROCEDURE:
1.
Collective — Adjust to maintain 85 to
107% NR.
2.
AIRSPEED — Adjust as desired.
RESTART − AUTOMATIC MODE
l PROCEDURE (NO RESTART FAULT OR
FADEC MANUAL LIGHTS ILLUMINATED):
NOTE
1.
Collective — Adjust to maintain 85 to
107% NR.
2.
AIRSPEED — Adjust as desired.
NOTE
Minimum rate of descent airspeed of 55
KIAS and minimum NR will allow pilot
more time for air restart.
3.
FUEL VALVE switch — ON.
4.
Throttle — Cutoff.
5.
START switch — Hold to start
position (start will latch after throttle
is placed to idle).
Minimum rate of descent airspeed of 55
KIAS and minimum NR will allow pilot
more time for air restart.
3.
Throttle — Closed.
4.
FADEC MODE switch — MAN.
5.
FUEL VALVE switch — ON.
6.
START switch — Hold to start
position (starter will not latch).
7.
NG — 12%.
8.
Throttle — Slowly advance out of
cutoff and stop advancing throttle at
light off.
9.
MGT — Allow to peak.
6.
NG — Between 12% and 50%.
7.
Throttle — Idle.
10. Throttle — Increase fuel flow by
modulating throttle to maintain MGT
within limits.
8.
MGT — Monitor.
11. START switch — Release at 50% NG.
9.
Throttle — Advance smoothly to FLY
detent position.
12. Throttle — Advance smoothly and
modulate to 100% NP.
If restart is unsuccessful, abort start and
secure engine as follows:
10. Throttle — Closed.
If restart is unsuccessful, abort start and
secure engine as follows:
11. FUEL VALVE switch — OFF.
12. Accomplish autorotative
and landing.
descent
13. Throttle — Closed.
17 DEC 2002
3-5
BHT-407-FM-1
TC APPROVED
14. FUEL VALVE switch — OFF.
15. Accomplish autorotative
and landing.
descent
3.
Increase in NG.
4.
Increase in TRQ.
PROCEDURE:
3-3-C.
ENGINE UNDERSPEED
NO CAUTION/WARNING/ADVISORY LIGHTS
ILLUMINATED.
1.
Throttle — Retard.
2.
NG or NP — Attempt to stabilize with
throttle and collective.
3.
FADEC MODE switch — MAN.
4.
NR — Maintain 95 to 100% with
throttle and collective.
INDICATIONS:
1.
Decrease in NG.
2.
Subsequent decrease in NP.
3.
Possible decrease in NR.
4.
Decrease in TRQ.
PROCEDURE:
1.
Collective — Adjust as required to
maintain 85 to 107% NR.
2.
Throttle — Confirm in FLY detent
position.
3.
Throttle — Position throttle to the
approximate bezel position that
coincides with the gauge indicated
NG.
4.
FADEC MODE switch — MAN.
5.
NR — Maintain 95 to 100% with
throttle and collective.
6.
Land as soon as practical.
3-3-D.
ENGINE OVERSPEED
(NO CAUTION/WARNING/ADVISORY LIGHTS
ILLUMINATED)
CAUTION
IF UNABLE TO MAINTAIN NR, NP, NG,
OR MGT, PREPARE FOR A POWER
OFF LANDING BY LOWERING
COLLECTIVE AND SHUTTING DOWN
ENGINE.
3-3-E.
ENGINE COMPRESSOR STALL
INDICATIONS:
1.
Engine pops.
2.
High or erratic MGT.
3.
Decreasing or erratic NG or NP.
4.
TRQ oscillations.
PROCEDURE:
1.
Collective
—
Reduce
maintain slow cruise flight.
power,
2.
MGT and NG — Check for normal
indications.
3.
ENG ANTI ICE switch — ON.
4.
PART SEP switch (if installed) — ON.
INDICATIONS:
1.
2.
3-6
Increase in NR.
Increase in NP.
17 DEC 2002
TC APPROVED
5.
BHT-407-FM-1
HEATER switch (if installed) — ON.
NOTE
Severity of compressor stalls will
dictate if engine should be shut down
and treated as an engine failure. Violent
stalls can cause damage to engine and
drive system components, and must be
handled as an emergency condition.
Stalls of a less severe nature (one or
two low intensity pops) may permit
continued operation of engine at a
reduced power level, avoiding
condition that resulted in compressor
stall.
If pilot elects to continue flight:
6.
Collective — Increase slowly to
achieve desired power level.
7.
MGT and NG — Monitor for normal
response.
8.
Land as soon as practical.
If pilot elects to shut down engine:
9.
Enter autorotation.
PROCEDURE:
1.
Throttle — Closed.
2.
FUEL VALVE switch — OFF.
NOTE
Starter will remain engaged until MGT
decreases to 150°C and then
automatically disengage. Starter may
b e m a n ua lly e ng a ge d b y ho ld in g
STARTER switch forward.
3.
STARTER switch — Ensure starter is
motoring engine until MGT stabilizes
at normal temperature.
4.
Shut down helicopter.
3-3-G.
INDICATIONS:
1.
Engine oil pressure below minimum.
2.
Engine oil
abnormally.
12. Collective — Adjust as required to
maintain 85 to 107% NR.
13. Cyclic — Adjust as required to
maintain desired AIRSPEED.
14. Prepare for power-off landing.
3-3-F.
ENGINE HOT START/SHUTDOWN
INDICATIONS:
pressure
fluctuating
PROCEDURE:
1.
Engine oil pressure and temperature
— Monitor.
2.
Land as soon as practical.
10. Throttle — Closed.
11. FUEL VALVE switch — OFF.
ENGINE OIL PRESSURE LOW
OR FLUCTUATING
3-3-H.
ENGINE OIL TEMPERATURE
HIGH
INDICATIONS:
1.
Engine oil temperature increasing
above normal.
2.
Engine oil
maximum.
temperature
above
PROCEDURE:
1.
Excessive MGT.
2.
Visible smoke or fire.
Land as soon as practical.
17 DEC 2002
3-7
BHT-407-FM-1
3-3-J.
TC APPROVED
DRIVESHAFT FAILURE
2.
Collective — Adjust as required to
maintain 85 to 107% NR.
WARNING
NOTE
FAILURE OF MAIN DRIVESHAFT TO
TR AN S MIS SIO N W ILL R ES U LT IN
COMPLETE LOSS OF POWER TO MAIN
ROTOR. ALTHOUGH COCKPIT
INDICATIONS FOR A DRIVESHAFT
FAILURE ARE SIMILAR TO AN ENGINE
OVERSPEED, IT IS IMPERATIVE THAT
AUTOROTATIVE
FLIGHT
P R O C E D U R E S B E E S T A B L IS H E D
IMMEDIATELY. FAILURE TO REACT
IMMEDIATELY TO LOW RPM AUDIO,
RPM LIGHT AND NP/NR TACHOMETER
CAN RESULT IN LOSS OF CONTROL.
INDICATIONS:
1.
Left yaw
2.
Rapid decrease in NR
3.
Rapid increase in NP
4.
LOW RPM audio horn
5.
Illumination of RPM light
6.
Possible increase in noise level due
to overspeeding engine and
driveshaft breakage.
Minimum rate of descent airspeed is 55
KIAS.
Maximum
glide
distance
airspeed is 80 KIAS.
3.
NOTE
To maintain tail rotor effectiveness do
not shutdown engine.
4.
Landing — Complete autorotative
landing.
5.
Complete helicopter shutdown.
3-3-K.
FADEC FAILURE
NOTE
Takeoff power may not be available in
the MAN mode. Maximum continuous
power will be available for all ambient
conditions.
INDICATIONS
1.
FADEC fail audio activated.
2.
FA D E C FA I L
illuminated.
Engine overspeed trip system will
activate at 118.5% NP causing fuel flow
to go to minimum. After initial
overspeed, FADEC will adjust fuel flow
to maintain engine at 100% NP.
3.
FA D E C M A N U A L c a u ti o n li g h t
illuminated.
4.
A U TO R E L I G H T a d v i s o r y li g h t
illuminated.
PROCEDURE:
5.
FADEC MODE switch MAN light
illuminated.
NOTE
1.
3-8
Cyclic — Adjust to obtain desired
autorotative airspeed.
Maintain
control.
heading
17 DEC 2002
and
attitude
warning
light
TC APPROVED
BHT-407-FM-1
PROCEDURE:
WARNING
WITHIN 2 TO 7 SECONDS AFTER
THE FADEC FAIL WARNING NR/NP
MAY
INCREASE
RAPIDLY,
REQUIRING POSITIVE MOVEMENTS
OF COLLECTIVE AND THROTTLE
TO CONTROL NR.
1.
2.
3.
Throttle — If time permits, match
throttle bezel position to NG
indication.
NR/NP — Maintain 95 to 100% with
collective and throttle.
FADEC MODE switch — Depress one
time, muting FADEC fail audio.
2.
FUEL VALVE switch — OFF.
3.
GEN switch — OFF.
4.
BATT switch — OFF.
5.
Rotor brake (if installed) — Engage.
6.
Exit helicopter.
3-4-B.
INDICATIONS:
1.
Smoke
2.
Fumes
3.
Fire
PROCEDURE:
1.
Inflight
—
autorotation.
2.
Throttle — Closed.
3.
FUEL VALVE switch — OFF.
4.
If time permits, FUEL BOOST/XFR
circuit breaker switches — OFF.
5.
Execute autorotative descent and
landing.
6.
BATT switch — OFF.
NOTE
Depressing FADEC MODE switch
one time, will only mute FADEC fail
au dio. T his ste p s hou ld n ot be
accomplished until pilot is firmly
established in MAN control.
4.
Land as soon as practical.
5.
Normal shutdown if possible.
3-4.
3-4-A.
FIRE
ENGINE FIRE DURING FLIGHT
ENGINE FIRE ON GROUND
INDICATIONS:
1.
Smoke
2.
Fumes
3.
Fire
Immediately
enter
NOTE
Do not restart engine until corrective
maintenance has been performed.
3-4-C.
CABIN SMOKE OR FUMES
INDICATIONS:
1.
Smoke
2.
Fumes
PROCEDURE:
1.
Throttle — Closed.
30 JUL 2008
Rev. 7
3-9
BHT-407-FM-1
TC APPROVED
PROCEDURE:
1.
Inflight — Start descent
2.
AIR COND BLO switch (if installed)
— OFF
3.
HEATER switch (if installed) — OFF
4.
All vents — Open
5.
Side windows — Open
If time and altitude permits:
6.
Source — Attempt to identify and
secure.
7.
If source is identified and smoke
and/or fumes still persist — Land as
soon as possible.
8.
3-5.
If source is identified and smoke
and/or fumes are cleared — Land as
soon as practical.
TAIL ROTOR
There is no single emergency procedure for
all types of antitorque malfunctions. One key
to a pilot successfully handling a tail rotor
emergency lies in the ability to quickly
recognize the type of malfunction that has
occurred.
3-5-A.
NOTE
Severity of initial reaction of helicopter
will be affected by AIRSPEED, CG,
power being used, and HD.
PROCEDURE:
3-5-A-1.
HOVERING
C l o s e t h r o tt l e a n d p e r f o r m a h o v e r i n g
autorotation landing. A slight rotation can be
expected on touchdown.
3-5-A-2.
IN-FLIGHT
Reduce throttle to idle, immediately enter
a u t o r o ta t i o n , a n d m a i n ta i n a m i n i m u m
AIRSPEED of 55 KIAS during descent.
NOTE
When a suitable landing site is not
available, vertical fin may permit
controlled flight at low power levels
and sufficient AIRSPEED. During final
stages of approach, a mild flare should
be executed, making sure all power to
rotor is off. Maintain helicopter in a
slight flare and smoothly use collective
to execute a soft, slightly nose-high
landing. Landing on aft portion of skids
will tend to correct side drift. This
technique will, in most cases, result in
a run-on type landing.
COMPLETE LOSS OF TAIL
ROTOR THRUST
CAUTION
This is a situation involving a break in drive
system (e.g., severed driveshaft), wherein tail
rotor stops turning and delivers no thrust.
INDICATIONS:
3-10
IN A RUN-ON TYPE LANDING AFTER
TOUCHING DOWN, DO NOT USE
CYCLIC TO REDUCE FORWARD
SPEED.
1.
Uncontrollable yawing to right (left
side slip).
3-5-B.
2.
Nose down tucking.
3.
Possible roll of fuselage.
This is a situation involving inability to
change tail rotor thrust (blade angle) with antitorque pedals.
17 DEC 2002
FIXED PITCH FAILURES
TC APPROVED
BHT-407-FM-1
INDICATIONS:
1.
Lack of directional response.
2.
Locked pedals.
approach with sufficiently low AIRSPEED
(zero if necessary) to attain a rate of descent
with a comfortable sideslip angle. (A decrease
i n N P d e c r e a s e s ta i l r o t o r t h r u s t . ) A s
collective is increased just before touchdown,
left yaw will be reduced.
NOTE
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. If helicopter is in a trimmed
condition when malfunction occurs,
TRQ and AIRSPEED should be noted
and helicopter flown to a suitable
landing area. Certain combinations of
TRQ, NR, and AIRSPEED will correct a
yaw attitude, and these combinations
should be used to land helicopter.
NOTE
Pull pedal stop emergency release to
ensure pedal stop is retracted.
HOVERING
Do not close throttle unless a severe right
yaw occurs. If pedals lock in any position at a
h o v e r, l a n d i n g f r o m a h o v e r c a n b e
accomplished with greater safety under
power-controlled flight rather than by closing
throttle and entering autorotation.
3-5-B-2.
IN-FLIGHT — LEFT PEDAL
APPLIED
In a high power condition, helicopter will yaw
to left when power is reduced. Power and
AIRSPEED should be adjusted to a value
w h e r e a c o m fo r ta b l e y a w a n g l e c a n b e
maintained. If AIRSPEED is increased, vertical
fin will become more effective and an
increased left yaw attitude will develop. To
accomplish landing, establish a power-on
IN-FLIGHT — RIGHT PEDAL
APPLIED
In cruise flight or reduced power situation,
helicopter will yaw to right when power is
increased. A low power, run-on type landing
will be necessary by gradually reducing
throttle to maintain heading while adding
collective to cushion landing. If right yaw
becomes excessive, close throttle completely.
3-6.
3-6-A.
PROCEDURE:
3-5-B-1.
3-5-B-3.
HYDRAULIC SYSTEM
LOSS OF HYDRAULIC
PRESSURE
INDICATIONS:
1.
HYDRAULIC SYSTEM caution light
illuminated.
2.
Grinding or howling noise from
pump.
3.
Increase in force required to move
flight controls.
4.
Feedback forces may be evident
during flight control movement.
PROCEDURE:
1.
Reduce AIRSPEED to 70 to 100
KIAS.
2.
HYD SYSTEM circuit breaker — Out.
If hydraulic power is not restored,
push breaker in.
3.
HYD SYS switch — HYD SYS; OFF if
hydraulic power is not restored.
17 DEC 2002
3-11
BHT-407-FM-1
TC APPROVED
4.
For extended flight set comfortable
AIRSPEED, up to 120 KIAS, to
minimize control forces.
5.
Land as soon as practical.
6.
A run-on landing at effective
translational
l i ft
speed
( a p p r o x i m a t e l y 1 5 k n o ts ) i s
recommended.
4.
3-7.
3-7-A.
3-6-B.
FLIGHT CONTROL ACTUATOR
MALFUNCTION
An actuator hardover can occur in any flight
control axis, but a cyclic cam jam will only
occur in the fore and aft axis. An actuator
hardover is manifested by uncommanded
movements of one or two flight controls. If
two controls move, the pilot will find one of
these controls will require a higher than
normal control force to oppose the
movement. This force cannot be “trimmed” to
zero without turning the HYD SYS switch OFF.
Once the hydraulic boost is OFF, the forces
on the affected flight control will be similar to
the “normal” hydraulic off forces.
Land as soon as possible using
procedure from paragraph 3-6-A
ELECTRICAL SYSTEM
GENERATOR FAILURE
INDICATIONS:
1.
GEN FAIL caution light illuminated.
2.
AMPS indicates 0.
3.
Voltmeter — Approximately 24 volts
PROCEDURE:
1.
GENERATOR
FIELD
G E N E R AT O R R E S E T
breakers — Check in.
and
circuit
2.
GEN switch — RESET; then GEN.
3.
If power is not restored, place GEN
switch to OFF; land as s oon as
practical.
NOTE
INDICATIONS:
1.
Uncommanded
movements
flight
control
2.
High flight control forces to oppose
movement in one axis
3.
Feedback forces only in affected
flight control axis
4.
Flight control forces
unaffected axis
normal
in
PROCEDURE:
3-12
1.
Attitude — Maintain
2.
HYD SYS switch — OFF
3.
AIRSPEED — Set to 70 to 100 KIAS
17 DEC 2002
With generator OFF, a fully charged
battery will provide approximately 21
minutes of power for basic helicopter
and one VHF COMM radio (35 minutes
with optional 28 ampere/hour battery).
3-7-B.
EXCESSIVE ELECTRICAL LOAD
INDICATIONS:
1.
AMPS indicates excessive load.
2.
Smoke or fumes.
PROCEDURE:
1.
GEN switch — OFF.
2.
BATT switch — OFF.
TC APPROVED
3.
BHT-407-FM-1
FUEL BOOST/XFR LEFT circuit
breaker switch — LEFT (on).
WARNING
PRIOR TO BATTERY DEPLETION,
ALTITUDE MUST BE REDUCED
BELOW 8000 FEET HP (JET A) OR 4000
FEET H P (JET B). UNUSABLE FUEL
MAY BE AS HIGH AS 150 POUNDS
AFTER THE BATTERY IS DEPLETED
DUE TO INABILITY TO TRANSFER
FUEL FROM FORWARD CELLS.
NOTE
With battery and generator OFF, an
80% charged battery will operate left
fuel boost pump and left fuel transfer
pump for approximately 1.7 hours (2.8
hours with optional 28 ampere/hour
battery).
4.
Airspeed — 60 KIAS or less.
NOTE
Pedal stop disengages with loss of
electrical power.
5.
Land as soon as practical.
NOTE
3.
Fuel will stop transferring from
f o r w a r d t o a ft f u e l c e l l a t
approximately 345 pounds total
indicated fuel.
PROCEDURE:
1.
LEFT and RIGHT FUEL BOOST/XFR
circuit breaker switches — Check
ON.
2.
Determine FUEL QTY in forward cell.
3.
Subtract quantity of fuel trapped in
forward cell from total to determine
usable fuel remaining.
4.
Plan landing accordingly.
3-9.
1.
High (approximately 15 pounds) fore
and aft cyclic control forces.
2.
Normal pedal, collective and lateral
cyclic control forces.
PROCEDURE:
1.
INDICATIONS:
L/FUEL XFR and R/FUEL
caution lights illuminate.
CYCLIC CAM JAM
A cyclic cam jam can only occur in the fore
and aft axis, whereas, an actuator hardover
can occur in any flight control axis. A cyclic
cam jam is manifested when a commanded
control movement requires a higher than
normal fore and aft spring force. The force felt
when moving the cyclic fore and aft with a
cam jam is the result of overriding a spring
capsule.
FUEL SYSTEM
DUAL FUEL TRANSFER FAILURE
1.
Last 150 pounds of fuel in forward
cell may not be usable.
INDICATIONS:
When throttle is repositioned to the idle
stop (during engine shutdown) the
PMA will go offline and the engine may
flame out.
3-8.
2.
XFR
Helicopter pitch attitude — Maintain
normal pitch attitudes with forward
or aft cyclic force.
19 FEB 2007
Rev. 5
3-13
BHT-407-FM-1
TC APPROVED
CAUTION
DO NOT TURN HYDRAULIC BOOST
OFF
2.
Land as soon as practical.
3-10. WARNING, CAUTION, AND
ADVISORY LIGHTS/
MESSAGES
Red warning lights/messages, fault
c o n d i t i o n s , a n d c o r r e c t iv e a c t io n s a r e
presented in Table 3-1.
Amber caution and White advisory lights/
messages and corrective actions are
presented in Table 3-2.
3-14
17 DEC 2002
TC APPROVED
BHT-407-FM-1
Table 3-1. Warning (red) lights
PANEL
WORDING
FAULT CONDITION
CORRECTIVE ACTION
BATTERY HOT
Battery overheating.
Turn BATT switch OFF and land as
soon as practical. If BATTERY RLY
light illuminates, turn GEN switch
OFF if conditions permit. Land as
soon as possible.
ENGINE OUT
NG less than 55 ± 1% and/or
FADEC senses ENGINE OUT.
Verify engine condition. Accomplish
engine failure procedure.
ENGINE OVSPD
NG greater than 110% or NP
versus TORQUE is above
maximum continuous limit
(102.4% NP at 100% TORQUE to
108.6% NP at 0% TORQUE).
Adjust throttle and collective as
necessary. Determine if engine is
controllable, if not shut down.
Maintenance action required before
next flight.
FADEC FAIL (During
start)
FADEC has detected a serious
malfunction.
Close throttle immediately. Engage
starter to reduce MGT. Applicable
maintenance action required prior to
next flight.
FADEC FAIL (Inflight)
FADEC has detected a
malfunction and an overspeed
may occur 2 to 7 seconds
following activation of FADEC
fail horn and illumination of
FADEC FAIL warning light.
Engine may underspeed
significantly prior to overspeed.
Any other FADEC related lights
may be illuminated.
Accomplish FADEC FAILURE
procedure, paragraph 3-3-K.
Applicable maintenance action
required prior to next flight.
RPM (with low RPM
audio)
NR below 95%.
Reduce collective and ensure
throttle is in FLY detent position.
Light will extinguish and audio will
cease when NR increases above
95%.
RPM (without audio)
NR above 107%.
Increase collective and/or reduce
severity of maneuver. Light will
extinguish when NR decreases
below 107%.
XMSN OIL PRESS
Transmission oil pressure is
below minimum.
Reduce power; verify fault with
gage. Land as soon as possible.
XMSN OIL TEMP
Transmission oil temperature is Reduce power; verify fault with
at or above red line.
gage. Land as soon as practical.
17 DEC 2002
3-15
BHT-407-FM-1
TC APPROVED
Table 3-2. Caution (amber) and advisory (white/green) lights
PANEL
WORDING
FAULT CONDITION
CORRECTIVE ACTION
AUTO RELIGHT (white)
Engine igniter is operating.
None.
NOTE
AUTO RELIGHT light will be
illuminated when ignition system is
activated.Ignition
system
is
activated:
1 - during start sequence
2 - in MANUAL mode with
above 55%
NG
3 - with FADEC detection of engine
out condition with NG above 50%.
BAGGAGE DOOR
Baggage compartment door not
securely latched.
BATTERY RLY
Battery relay has malfunctioned If BATTERY HOT light is illuminated,
to closed (ON) position with
turn GEN switch OFF if conditions
BATT switch OFF. Battery is still permit. Land as soon as possible.
connected to DC BUSS.
CHECK INSTR
TRQ, MGT, or NG is about to or
has detected an exceedance.
Flashing LCD trend arc and
digital display indicates
impending exceedance. Letter E
in digital display indicates an
exceedance has occurred.
Reduce engine power if possible.
Press INSTR CHK button to display
magnitude of exceedance. Refer to
BHT-407-MD-1.
CYCLIC CENTERING
Cyclic stick is not centered.
Reposition cyclic stick to center
position to extinguish CYCLIC
CENTERING light.
ENGAGED
Information system status.
None.
ENGINE ANTI-ICE
(white)
ANTI-ICE switch ON. Engine
receiving anti-icing air.
If light (if installed) remains
illuminated with ENGINE ANTI-ICE
switch OFF, avoid operations
requiring maximum power.
ENGINE CHIP
Ferrous particles in engine oil.
Land as soon as possible.
3-16
17 DEC 2002
Close door securely before flight. If
light illuminates during flight, land
as soon as practical.
TC APPROVED
BHT-407-FM-1
Table 3-2. Caution (amber) and advisory (white/green) lights (Cont)
PANEL
WORDING
FADEC DEGRADED
(Inflight)
FAULT CONDITION
CORRECTIVE ACTION
FADEC ECU operation is
degraded which may result in NR
droop, NR lag, or reduced
maximum power capability.
Remain in AUTO mode. Fly
helicopter smoothly and
nonaggressively. Land as soon as
practical.
NOTE
It may be necessary to use FUEL
VALVE switch to shut down engine
after landing.
Applicable maintenance action
required prior to next flight.
FADEC DEGRADED
FADEC ECU has recorded a fault Position throttle to idle; if light
(With engine shutdown) during previous flight or a
extinguishes, fault is from previous
current fault has been detected. flight. Applicable maintenance
action required prior to next flight.
FADEC FAULT
PMA and or MGT, NP or NG
automatic limiting circuit(s) not
functional.
Remain in AUTO mode. Land as
soon as practical. Applicable
maintenance action required prior
to next flight.
FADEC MANUAL
FADEC is operating in MANUAL
mode. No automatic governing is
available. AUTO RELIGHT light
will be illuminated.
Fly helicopter smoothly and
nonaggressively. Maintain NR with
coordinated throttle and collective
movements. Land as soon as
practical.
FLOAT ARM
FLOAT ARM switch is ON. Float
inflation solenoid is armed.
Normal operation for takeoff and
landing over water. FLOAT ARM
switch — OFF. If light remains
illuminated, FLOATS circuit breaker
— Out. Land as soon as practical.
NOTE
With float inflation solenoid armed,
flight should not exceed 60 KIAS and
500 feet AGL.
FLOAT TEST (green)
Float system in test mode.
None.
17 DEC 2002
3-17
BHT-407-FM-1
TC APPROVED
Table 3-2. Caution (amber) and advisory (white/green) lights (Cont)
PANEL
WORDING
FAULT CONDITION
CORRECTIVE ACTION
FUEL FILTER
Airframe fuel filter in impending
bypass.
Land as soon as practical. Clean
before next flight.
FUEL LOW
100 ±10 pounds of fuel remain in Verify FUEL QTY. Land as soon as
aft tank.
practical.
R/FUEL BOOST
Right fuel boost pump has failed. If practical, descend below 8000 feet
HP if fuel is Jet A or 4000 feet HP if
fuel is Jet B to prevent fuel
starvation if other fuel boost pump
fails or has low output pressure.
Land as soon as practical.
WARNING
IF BOTH FUEL BOOST PUMPS FAIL,
ALTITUDE MUST BE REDUCED TO
BELOW 8000 FEET HP (JET A) OR
4000 FEET Hp (JET B). LAND AS
SOON AS POSSIBLE.
L/FUEL BOOST
Left fuel boost pump has failed.
FUEL VALVE
Fuel valve position differs from Check FUEL VALVE circuit breaker
FUEL VALVE switch indication or in. Land a soon as practical. If on
FUEL VALVE circuit breaker out. ground, cycle FUEL VALVE switch.
L/FUEL XFR
Left fuel transfer pump has
failed.
3-18
Rev. 5
19 FEB 2007
If practical, descend below 8000 feet
HP if fuel is Jet A or 4000 feet HP if
fuel is Jet B to prevent fuel
starvation if other fuel boost pump
fails or has low output pressure.
Land as soon as practical.
Land as soon as practical.
TC APPROVED
BHT-407-FM-1
Table 3-2. Caution (amber) and advisory (white/green) lights (Cont)
PANEL
WORDING
FAULT CONDITION
CORRECTIVE ACTION
CAUTION
IF BOTH FUEL TRANSFER PUMPS
FAIL, UNUSABLE FUEL MAY BE AS
HIGH AS 150 POUNDS DUE TO
INABILITY TO TRANSFER FUEL
FROM FORWARD CELL. LAND AS
SOON AS PRACTICAL.
NOTE
Under normal fuel transfer
conditions, helicopters S/N 53000
through 53174 L/FUEL XFR and
R/FUEL XFR lights will illuminate
for 2.5 minutes and then
extinguish.
This
indicates
transfer is complete and transfer
pumps have been automatically
turned off. Helicopters S/N 53175
and
subsequent
inhibit
illumination of the lights.
R/FUEL XFR
Right fuel transfer pump has
failed.
Land as soon as practical.
GEN FAIL
Generator not connected to DC
BUSS.
Verify fault with AMPS gauge. GEN
switch — RESET, then ON. If GEN
FAIL light remains illuminated, GEN
switch — OFF. Land as soon as
practical.
HEATER OVERTEMP
An overtemp condition has been Turn HEATER switch OFF
detected by a temperature probe immediately.
either under pilot seat, copilot
seat, or in vertical tunnel.
19 FEB 2007
Rev. 5
3-19
BHT-407-FM-1
TC APPROVED
Table 3-2. Caution (amber) and advisory (white/green) lights (Cont)
PANEL
WORDING
FAULT CONDITION
CORRECTIVE ACTION
HYDRAULIC SYSTEM
Hydraulic pressure below limit.
Verify HYD SYS switch position.
Accomplish hydraulic system
failure procedure (refer to paragraph
3-6).
LITTER DOOR
Litter door not securely latched. Close door securely before flight. If
light illuminates during flight, land
as soon as practical.
PEDAL STOP
Pedal Restrictor Control Unit has VNE — 60 KIAS.
detected a failure of part of
system.
PEDAL STOP emergency release —
Pull.
Land as soon as practical.
RESTART FAULT
(white)
FADEC ECU has detected a fault Remain in AUTO mode. Plan landing
which will not allow engine to be site accordingly.
restarted in AUTO mode.
Applicable maintenance action
required prior to next flight.
NOTE
When throttle is repositioned to idle
stop (during engine shutdown) the
PMA will go offline and engine may
flameout.
START
(white)
Start relay is in START mode.
If START switch has not been
engaged and there is zero indication
on AMPS gage; START relay has
malfunctioned and helicopter is on
battery power. START circuit
breaker — Out. Land as soon as
practical.
T/R CHIP
Ferrous particles in tail rotor
gearbox oil.
Land as soon as possible.
XMSN CHIP
Ferrous particles in transmission Land as soon as possible.
oil.
3-20
17 DEC 2002
TC APPROVED
BHT-407-FM-1
Section 4
PERFORMANCE
4
TABLE OF CONTENTS
Page
Number
Subject
Paragraph
INTRODUCTION .........................................................................
POWER ASSURANCE CHECK..................................................
DENSITY ALTITUDE...................................................................
HEIGHT – VELOCITY ENVELOPE.............................................
HOVER CEILING.........................................................................
NOT USED ..................................................................................
CLIMB AND DESCENT...............................................................
CLIMB ...................................................................................
AUTOROTATION .................................................................
AIRSPEED CALIBRATION.........................................................
NOT USED ..................................................................................
NOISE LEVELS...........................................................................
FAR PART 36 STAGE 2 NOISE LEVEL.................................
CANADIAN AIRWORTHINESS MANUAL CHAPTER
516 AND ICAO ANNEX 16 NOISE LEVEL.............................
4-1....................
4-2....................
4-3....................
4-4....................
4-5....................
4-6....................
4-7....................
4-7-A................
4-7-B................
4-8....................
4-9....................
4-10..................
4-10-A..............
4-3
4-3
4-4
4-4
4-4
4-5
4-5
4-5
4-6
4-6
4-6
4-6
4-6
4-10-B..............
4-6
LIST OF FIGURES
Title
Power assurance check ............................................................
Density altitude ..........................................................................
Altitude vs gross weight for height – velocity diagram .........
Height – velocity diagram .........................................................
Hover ceiling wind accountability chart –
below 14,000 feet HD ................................................................
Hover ceiling wind accountability chart –
between 14,000 and 17,000 feet HD ........................................
Hover ceiling wind accountability chart –
above 17,000 feet HD ...............................................................
Hover ceiling IGE .......................................................................
Hover ceiling OGE .....................................................................
Rate of climb – takeoff power ...................................................
Rate of climb – maximum continuous power..........................
Autorotation glide distance ......................................................
Airspeed installation correction ...............................................
Figure
Number
Page
Number
4-1....................
4-2....................
4-3....................
4-4...................
4-7
4-8
4-9
4-10
4-5...................
4-11
4-5A ................
4-12
4-5B ................
4-6...................
4-7...................
4-8...................
4-9...................
4-10.................
4-11.................
4-13
4-14
4-22
4-30
4-40
4-50
4-51
17 DEC 2002
4-1/4-2
TC APPROVED
BHT-407-FM-1
Section 4
PERFORMANCE
4
4-1.
INTRODUCTION
SOLUTION:
Performance data presented herein are
derived from engine manufacture's
specification power for engine less
installation losses. These data are applicable
to basic helicopter without any optional
equipment that would appreciably affect lift,
drag, or power available.
E n t e r P o w e r a s s u ra n c e c h e c k c h a r t a t
observed TORQUE (70%), proceed vertically
down to intersect H P (6000 feet), follow
horizontally to intersect indicated OAT (10°C),
then drop vertically to read maximum
allowable MGT.
4-2.
If actual MGT is less than or equal to chart
MGT, engine performance equals or exceeds
minimum specification and performance data
contained in this manual can be achieved.
POWER ASSURANCE CHECK
A Power assurance check chart (Figure 4-1) is
provided for Rolls-Royce model 250-C47B
engine. This chart indicates maximum
a l l o w a b l e M GT f o r a n e n g i n e m e e t i n g
minimum Rolls-Royce specification. Engine
must develop required torque without
e x c e e d i n g c h a r t M GT i n o r d e r t o m e e t
performance data contained in this manual.
Figure 4-1 may be used to periodically
monitor engine performance.
If actual MGT is greater than chart MGT,
engine performance is less than minimum
s pe c ific at io n a nd al l pe rfo r man c e d ata
contained in this manual cannot be achieved.
Refer to appropriate maintenance manual to
determine cause of low power (high MGT).
NOTE
To perform power assurance check, turn off
all sources of bleed air, including ENGINE
AN TI-IC ING. Es tab lish leve l flig ht at an
AIRSPEED of 85 to 105 KIAS or VNE,
wh ic heve r is low er. C hec k may also be
c o n d u c t e d i n a h o v e r p r i o r t o ta k e o f f ,
depending on ambient conditions and gross
weight.
Record following information from cockpit
instruments:
EXAMPLE:
HP
6000 feet
OAT
10°C
MGT
Actual reading
TORQUE
70%
Chart may also be used to determine
minimum sp ecificatio n pow er for
actual MGT. Using above example,
enter chart at actual MGT (675°C,
proceed up to OAT (10°C), across to HP
(6000 feet), and up to read minimum
torque available (70%). If actual power
is equal to or greater than chart torque,
engine performance equals or exceeds
minimum
specification
and
performance data contained in this
manual can be achieved. If actual
torque indication is less than chart
torque, engine performance is less
than minimum specification and all
performance in this manual cannot be
achieved. Refer to appropriate
maintenance manual to determine
cause of low power.
17 DEC 2002
4-3
BHT-407-FM-1
4-3.
DENSITY ALTITUDE
A Density altitude and temperature
conversion chart (Figure 4-2) is provided to
aid in calculation of performance and
limitations. HD is an expression of density of
air in terms of height above sea level; hence,
the less dense the air, the higher the HD. For
standard conditions of temperature and
pressure, H D is same as H P. As temperature
increases above standard for an altitude, HD
will also increase to values higher than H P.
Figure 4-2 expresses HD as a function of HP
and temperature.
Density altitude chart also includes the
inverse of the square root of the density ratio
(1 / σ ), w h ic h i s u se d to c a lc ula te tru e
airspeed by the following relation:
KTAS = KCAS×1/ σ
TC APPROVED
(Figure 4-4) is valid only when helicopter
gross weight does not exceed limits of the
Altitude vs Gross Weight for Height – Velocity
diagram (Figure 4-3). Four envelopes (Gross
Weight Regions) are specified. Each Gross
Weight Region applies for all gross weights
within its boundaries. No interpolation is
allowed.
For a given ambient outside air temperature,
pressure altitude, and gross weight, the
appropriate limiting envelope (Region A, B, C,
or D) can be determined. Using Figure 4-3
(Altitude VS Gross Weight), move upward
vertically from entry OAT to pressure altitude.
From that point, move right horizontally to
determine the correct weight region.
(Examples: 15°C at Sea Level at 5000 pounds
GW = Region B, and 30°C at 2000 feet
pressure altitude at 5 000 pounds G W =
Region D) Once the correct weight region has
been determined (A, B, C, or D), the
corresponding Avoid area is selected from
Figure 4-4 (Height – Velocity diagram).
EXAMPLE:
4-5.
HOVER CEILING
If ambient temperature is -15°C and HP is 7000
feet, find HD, 1/ σ , and true airspeed for 100
KCAS.
SOLUTION:
Enter bottom of chart at -15°C.
Move vertically upward to 7000 feet HP
line.
From this point, move horizontally to
left and read H D of 5000 feet, move
horizontally to right and read 1/ σ =
1.08.
True airspeed = KCAS × 1/ σ = 100 ×
1.08 = 108 KTAS.
4-4.
HEIGHT – VELOCITY
ENVELOPE
T h e H e i g h t – Ve lo c i t y e n v e l o p e c h a r ts
(Figures 4-3 and 4-4) define conditions from
which a safe landing can be made on a
smooth, level, firm surface; following an
engine failure. The Height – Velocity diagram
4-4
17 DEC 2002
NOTE
Hover performance charts are based
on 100% ROTOR RPM.
Satisfactory stability and control have been
demonstrated in each area of the Hover
ceiling charts with winds as depicted on the
Hover ceiling wind accountability chart
(Figures 4-5, 4-5A and 4-5B).
Hover ceiling – in ground effect charts (Figure
4-6) and Hover ceiling – out of ground effect
charts (Figure 4-7) present hover performance
as allowable gross weight for conditions of HP
a n d O AT. T h e s e h o v e r i n g w e i g h ts a r e
obtainable in zero wind conditions. Each
chart is divided into two areas: Area A (non
shaded area) and Area B (shaded area).
For the data presented below 14,000 ft H D ,
Area A of the hover ceiling charts presents
hov er pe rfo rma nce (relativ e to GW) for
conditions where adequate control margins
exist for all relative wind conditions up to 35
knots for lateral CG not exceeding ±2.5 inches
(±63 mm); and up to 17 knots, for lateral CG
not exceeding ±4.0 inches (±102 mm); for
TC APPROVED
hover, takeoff and landing. Area B of the
h o v e r c e i l i n g c h a r ts p r e s e n ts h o v e r
performance (relative to GW) for conditions
where adequate control margins exist for
relative winds within ±45° of the nose of
helicopter up to 35 knots for lateral CG not
exceeding ±2.5 inches (±63 mm), and up to 17
knots for lateral CG not exceeding ±4.0 inches
(±102 mm); for hover, takeoff and landing.
For data presented between 14,000 and 17,000
ft H D , Area A of the hover ceiling charts
presents hover performance (relative to GW)
fo r c o n d i ti o n s w h e r e a d e q u a te c o n t ro l
margins exist for all relative wind conditions
up to 20 knots for lateral CG not exceeding
±2.5 inches (±63 mm); for hover, takeoff and
landing. Area B of the hover ceiling charts
presents recommended azimuth for takeoff
and landing for all relative winds within ±30°
of nose of helicopter for lateral CG not
exceeding ±2.5 inches (±63 mm).
For data presented above 17,000 ft HD, there
i s n o A r e a A . A r e a B p r e s e n ts h o v e r
performance (relative to GW) for conditions
where adequate control margins exist for all
relative winds within ±30° of the nose of the
helicopter for lateral CG not exceeding ±2.5
inches (±63 mm); for hover, takeoff and
landing
The following example uses a Hover ceiling
chart at takeoff power. The example is typical
for use with all other Hover ceiling charts.
EXAMPLE:
What IGE GW hover capability could be
expected for the following conditions:
A.
B.
C.
D.
HEATER and ANTI ICE – OFF
HP – 6000 feet
OAT – +20°C
TAKE OFF POWER
BHT-407-FM-1
SOLUTION:
Use Hover ceiling IGE – takeoff power chart
(sheet 1 of Figure 4-6).
A.
B.
C.
D.
Enter OAT scale at +20 °C.
Move upward to 6000 feet HP curve.
Move horizontally to +20 °C curve.
Drop down to read maximum
external gross weight of 5400
pounds (IGE hover capability
exceeds maximum internal GW of
5000 pounds).
4-6.
NOT USED
4-7.
CLIMB AND DESCENT
4-7-A.
CLIMB
Rate of climb charts are presented for various
combinations of power settings and ENGINE
ANTI-ICING switch positions. Refer to Figures
4-8 and 4-9.
Recommended best rate of climb airspeed is
60 KIAS.
Reduce rate of climb data 100 feet per minute
when operating with any combination of
door(s) removed.
The following example uses a Rate of climb
chart at takeoff power. The example is typical
for use with all other Rate of climb charts.
EXAMPLE:
Find the maximum rate of climb that can be
attain ed u sing take off pow er unde r the
following conditions:
HEATER
OFF
ENGINE ANTI-ICING
OFF
OAT
10°C
HP
14,000 feet
GW
3500 pounds
17 DEC 2002
4-5
BHT-407-FM-1
TC APPROVED
SOLUTION:
Enter appropriate gross weight chart (sheet 3
of Figure 4-8). At H P scale of 14000 feet
proceed horizontally to temperature of 10°C.
Drop down vertically and read a rate of climb
of 1700 feet per minute.
4-7-B.
AUTOROTATION
Refer to Figure 4-10 for autorotational glide
distance as a function of altitude.
4-8.
AIRSPEED CALIBRATION
Refer to Figure 4-11 for airspeed installation
correction during level flight and climb.
4-9.
NOT USED
4-10. NOISE LEVELS
4-10-A. FAR PART 36 STAGE 2
NOISE LEVEL
This aircraft is certified as a Stage 2 helicopter
as prescribed in FAR Part 36, Subpart H, for
g r o s s w e ig h ts u p t o a n d in c l u d i n g t h e
certificated maximum takeoff and landing
weight of 5000 pounds (2268 kilograms).
There are no operating limitations to meet any
of the noise requirements.
The following noise level complies with FAR
Part 36, Appendix J, Stage 2 noise level
requirements. It was obtained by analysis of
approved data from noise tests conducted
u n d e r t h e p r o v i s i o n s o f FA R P a r t 3 6 ,
Amendment 36-20.
levels of this aircraft are or should be
a c c e p ta b l e o r u n a c c e p t a b l e f o r
operations at, into, or out of any
airport.
V H is 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
(77°F) ambient conditions at the relevant
maximum certificated weight. The value of VH
thus defined for this aircraft is 127 KTAS.
4-10-B. CANADIAN AIRWORTHINESS
MANUAL CHAPTER 516 AND
ICAO ANNEX 16 NOISE LEVEL
This aircraft complies with the noise emission
standards applicable to the aircraft as set out
b y t h e I n t e r n a t i o n a l C i v i l Av i a t i o n
Organization (ICAO) in Annex 16, Volume 1,
Chapter 11, for gross weights up to and
including the certificated maximum takeoff
and landing weight of 5000 pounds (2268
kilograms). There are no operating limitations
to meet any of the noise requirements.
The following noise level complies with ICAO
Annex 16, Volume 1, Chapter 11 noise level
requirements. It was obtained by analysis of
approved data from noise tests conducted
under the provisions of ICAO Annex 16,
Volume 1, Third Edition-1993.
The flyover noise level for the Model 407 is
84.6 dBA SEL.
NOTE
The certified flyover noise level for the Model
407 is 85.1 dBA SEL.
NOTE
No determination has been made by
the certifying authorities that the noise
4-6
17 DEC 2002
ICAO Annex 16, Volume 1, Chapter 11
a pp r ov a l is ap p lic a b le o n ly aft er
endorsement by the Civil Aviation
Authority of the country of aircraft
registration.
TC APPROVED
BHT-407-FM-1
4
Figure 4-1. Power assurance check
17 DEC 2002
4-7
BHT-407-FM-1
24
TC APPROVED
,00
25
1.46
1.44
1.42
1.40
1.38
1.36
1.34
1.32
0
EXAMPLE: If OAT is -15°C and
HP is 7000 ft, HD is 5000 ft
22
and 1 /
σ is 1.08
20
0
,00
20
18
1.30
16
1.28
12
1.24
1.22
10
0
,00
10
8
P
E
UR
SS
E
R
DE
IT U
T
AL
-F
1.20
T
1.18
1.16
1.14
σ
0
,00
1.12
1/
DENSITY ALTITUDE - FT x 1000
15
AY
DD
AR
ND
STA
14
1.26
1.10
6
1.08
00
5,0
4
1.06
1.04
2
1.02
0
A
SE
-2
1.00
L
VE
LE
0.98
0.96
0
00
-5,
-4
0.94
0.92
-6
0
,00
-10
-8
-50
-40
-30
-20
-10
0
10
20
30
OUTSIDE AIR TEMPERATURE - °C
40
50
0.90
60
M407_FM-1__FIG_4-2.WMF
Figure 4-2. Density altitude
4-8
17 DEC 2002
TC APPROVED
BHT-407-FM-1
ALTITUDE VS GROSS WEIGHT
FOR HEIGHT-VELOCITY DIAGRAM
GROSS WEIGHT - KG x 100
13 14 15 16 17 18 19 20 21 22
14
,0
00
14,000 FT HD
00
60
S
OS
40
GH
EI
00
W
N
IO
20
00
B
EL
S
LE
V
A
-2
0
00
SE
A
T
-F
Hp
EG
-20
TR
-40
C
MAXIMUM INTERNAL GROSS WEIGHT
10
,
80
00
D
GR
MINIMUM OAT LIMIT
9000 FT HD
AT
UM O
00
0
12
MAXIM
,0
00
DEMONSTRATED
TO 9000 FEET
DENSITY ALTITUDE
0
20
OAT - °C
40
60
30
34
38
42
46
50
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-3.EMF
Figure 4-3. Altitude vs gross weight for height – velocity diagram
17 DEC 2002
4-9
BHT-407-FM-1
TC APPROVED
HEIGHT-VELOCITY DIAGRAM
FOR GROSS WEIGHT REGIONS A TO D
800
400
350
160
140
120
E
SKID HEIGHT - FEET
450
S
500
180
EA
AR
550
OID
600
200
AV
650
220
NOTE: LOW HOVER POINT IS
AT 6 FT SKID HEIGHT
100
300
SKID HEIGHT - METERS
700
E
LIN
LIN
E
IS
IS
E
LIN
TH
TH
W
LIN HIS
LO
IS
T
OW
BE
TH
EL
OW
W
D"
"B
EL
LO
N"
B
E
"A
"
B
B
GIO
"
C"
RE
ON
N"
ON
GI
GI
GIO
RE
RE
RE
750
240
80
250
200
60
AVOID
150
40
100
20
50
0
0
0
10
20
30
40
50
60
70
80
90
100
110
120
INDICATED AIRSPEED - KNOTS
M407_FM-1__FIG_4-4.EMF
Figure 4-4. Height – velocity diagram
4-10
17 DEC 2002
TC APPROVED
BHT-407-FM-1
SHADED AREA (AREA B)
PRESENTS HOVER PERFORMANCE
(RELATIVE TO GW) FOR CONDITIONS
WHERE ADEQUATE CONTROL MARGINS
EXIST FOR RELATIVE WINDS WITHIN ± 45°
OF NOSE OF HELICOPTER UP TO 35
KNOTS FOR LATERAL CG NOT EXCEEDING
± 2.5 INCHES (± 63 mm), AND UP TO 17
KNOTS FOR LATERAL CG NOT EXCEEDING
± 4.0 INCHES (± 102 mm); FOR HOVER,
TAKEOFF AND LANDING.
ALTITUDE BELOW
14000 FT HD
± 45°
ALL
AZIMUTHS
NON-SHADED AREA (AREA A)
PRESENTS HOVER PERFORMANCE
(RELATIVE TO GW) FOR CONDITIONS
W H E R E A D E Q U AT E C O N T R O L
MARGINS EXIST FOR ALL RELATIVE
WIND CONDITIONS UP TO 35 KNOTS
FOR LATERAL CG NOT EXCEEDING
± 2.5 INCHES (± 63 mm), AND UP TO
17 KNOTS FOR LATERAL CG NOT
EXCEEDING ± 4.0 INCHES (± 102 mm);
FOR HOVER, TAKEOFF AND LANDING.
M407_FM-1_FIG_4-5.WMF
Figure 4-5. Hover ceiling wind accountability chart – below 14,000 feet HD
17 DEC 2002
4-11
BHT-407-FM-1
TC APPROVED
ALTITUDE BETWEEN
14000 AND 17000 FT HD
SHADED AREA (AREA B)
PRESENTS RECOMMENDED AZIMUTH
FOR TAKEOFF AND LANDING FOR ALL
RELATIVE WINDS WITHIN ± 30° OF NOSE
OF HELICOPTER FOR LATERAL CG NOT
EXCEEDING ± 2.5 INCHES (± 63 mm).
± 30°
ALL
AZIMUTHS
NON-SHADED AREA (AREA A)
PRESENTS HOVER PERFORMANCE
(RELATIVE TO GW) FOR CONDITIONS
W H E R E A D E Q U AT E C O N T R O L
MARGINS EXIST FOR ALL RELATIVE
WIND CONDITIONS UP TO 20 KNOTS
FOR LATERAL CG NOT EXCEEDING
± 2.5 INCHES (± 63 mm); FOR HOVER,
TAKEOFF AND LANDING.
M407_FM-1__FIG_4-5A.WMF
Figure 4-5A. Hover ceiling wind accountability chart – between 14,000 and 17,000 feet HD
4-12
17 DEC 2002
TC APPROVED
SHADED AREA (AREA B)
PRESENTS HOVER PERFORMANCE
( RE L AT IV E TO GW ) F O R
CONDITIONS WHERE ADEQUATE
CONTROL MARGINS EXIST FOR ALL
WINDS DIRECTLY OFF THE NOSE OF
THE HELICOPTER; FOR HOVER,
TAKEOFF AND LANDING.
BHT-407-FM-1
ALTITUDE ABOVE
17000 FT HD
M407_FM-1_FIG_4-5B.WMF
Figure 4-5B. Hover ceiling wind accountability chart – above 17,000 feet HD
17 DEC 2002
4-13
BHT-407-FM-1
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
-3
16
,
00
0
,0
00
20
,0
0
12
00
,0
10
00
80
00
40
00
60
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
20
OAT - °C
40
60 32
36
40
44
48
.7
REFER TO
FIG. 4-5
51
00
50
0
0
-4
40
-20
°C
AT
T
30
OA
MINIMUM OAT
-
0
UM
IM
UM O
MAXIM
-40
AT
O
0
-1
0
0
-2
X
MA
14,000 FT HD
10
14
0
REFER TO
FIG. 4-5A
52
56
EXTERNAL = 6000 LB
18
,0
20
,0
00
20,000 FT HD
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(1_OF_8).WMF
Figure 4-6. Hover ceiling IGE (sheet 1 of 8)
4-14
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
-3
,0
00
REFER TO
FIG. 4-5A
12
00
,0
0
-2
00
°C
0
-4
10
-
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
5
AT
80
AT
0
UM O
MAXIM
MINIMUM OAT
O
0
-1
,0
0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(2_OF_8).WMF
Figure 4-6. Hover ceiling IGE (sheet 2 of 8)
17 DEC 2002
4-15
BHT-407-FM-1
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
AREA B
00
17,000 FT HD
16
,0
18
,0
00
20
,0
00
20,000 FT HD
00
14
,0
30
0 -4
REFER TO
FIG. 4-5A
14,000 FT HD
00
,0
12
00
,0
-
00
20
-2
00
0
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
LE
V
A
T
-F
P
H
-20
EL
20
00
40
00
60
00
AT
80
°C
10
AT
10
UM O
MINIMUM OAT
O
0 0
0 -1
-2
MAXIM
-40
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(3_OF_8).WMF
Figure 4-6. Hover ceiling IGE (sheet 3 of 8)
4-16
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
°C
00
,0
10
00
-
0
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
5
AT
80
AT
0
MINIMUM OAT
O
0 1
0 -2 -
UM O
MAXIM
12
,0
0
0 -3
-4
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(4_OF_8).WMF
Figure 4-6. Hover ceiling IGE (sheet 4 of 8)
17 DEC 2002
4-17
BHT-407-FM-1
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
AREA B
00
17,000 FT HD
16
,0
18
,0
00
20
,0
00
20,000 FT HD
14
0
30 -2
0 -4 -10 0
,0
00
REFER TO
FIG. 4-5A
00
,0
00
°C
-20
00
0
-2
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
EL
LE
V
REFER TO
FIG. 4-5
A
SE
MAX GW INTERNAL
= 5000 LB
00
20
T
-F
P
H
-40
50
51.7
40
00
60
40
00
30
80
-
10
AT
20
AT
T
OA
MINIMUM OAT
O
10
UM
UM O
XIM
MA
MAXIM
12
,0
00
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(5_OF_8).EMF
Figure 4-6. Hover ceiling IGE (sheet 5 of 8)
4-18
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
,0
00
0
14
-4
REFER TO
FIG. 4-5A
14,000 FT HD
0
,0
0
12
00
,0
10
°C
00
-
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
5
AT
80
AT
0
MINIMUM OAT
O
0 - 10
0 -2
-3
UM O
MAXIM
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(6_OF_8).WMF
Figure 4-6. Hover ceiling IGE (sheet 6 of 8)
17 DEC 2002
4-19
BHT-407-FM-1
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
,
00
0
17,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
0
12
,0
0
00
,0
10
00
°C
0
80
AT
O
0
0 20 -1
0 -3 -4
AT
UM O
MAXIM
MINIMUM OAT
AREA B
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
MAX GW
INTERNAL = 5000 LB
00
18
,0
20
,0
00
20,000 FT HD
60
00
10
-40
-20
0
-2
00
SE
P
H
-F
A
T
LE
VE
L
20
00
40
00
20
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(7_OF_8).EMF
Figure 4-6. Hover ceiling IGE (sheet 7 of 8)
4-20
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
-
12
00
,0
10
80
00
MINIMUM OAT
AT
00
°C
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-20
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
5
40
00
60
0
-40
O
AT
0
-1
UM O
MAXIM
20
0 30 -4 -
,0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-6_(8_OF_8).WMF
Figure 4-6. Hover ceiling IGE (sheet 8 of 8)
17 DEC 2002
4-21
BHT-407-FM-1
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
,0
0
0
0
-3
0
00
,0
60
0
-4
00
T
30
OA
00
°C
10
-
0
-1
20
M
AT
80
AT
U
XIM
MINIMUM OAT
O
0
-2
12
10
MA
UM O
MAXIM
-2
00
0
A
SE
0
20
OAT - °C
40
60 32
36
40
44
48
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
LE
VE
L
20
T
-F
P
H
-20
.7
51
-40
50
00
40
00
40
REFER TO
FIG. 4-5
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(1_OF_8).WMF
Figure 4-7. Hover ceiling OGE (sheet 1 of 8)
4-22
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
0
,0
0
12
00
,0
10
°C
0
-2
00
-
0
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
AT
80
AT
-3
MINIMUM OAT
O
0
-4
0 0 5
-1
UM O
MAXIM
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(2_OF_8).WMF
Figure 4-7. Hover ceiling OGE (sheet 2 of 8)
17 DEC 2002
4-23
BHT-407-FM-1
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
AREA B
00
17,000 FT HD
16
,0
18
,0
00
20
,0
00
20,000 FT HD
14
-4
,0
00
REFER TO
FIG. 4-5A
,0
12
00
°C
80
-
AT
-20
-2
00
0
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
LE
V
A
T
-F
P
H
-40
REFER TO
FIG. 4-5
EL
20
00
40
00
60
00
20
MINIMUM OAT
AT
UM O
10
,0
00
MAXIM
O
0 10 0 1 0
0 -2 -
00
0 -3
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(3_OF_8).WMF
Figure 4-7. Hover ceiling OGE (sheet 3 of 8)
4-24
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
12
,0
0
0
0
-4
00
,0
00
AT
O
0
-1
-
AT
80
0
-2
10
0
°C
MINIMUM OAT
-3
UM O
MAXIM
60
00
0
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
5
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(4_OF_8).WMF
Figure 4-7. Hover ceiling OGE (sheet 4 of 8)
17 DEC 2002
4-25
BHT-407-FM-1
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
14
00
,0
00
00
60
30
40
00
40
REFER TO
FIG. 4-5
00
0
-2
H
P
SE
-F
A
T
LE
V
EL
20
00
50
51.7
-20
°C
80
-
T
20
OA
AT
-40
AT
UM
10
XIM
UM O
MINIMUM OAT
MA
MAXIM
12
,0
00
14,000 FT HD
O
0 1 0 0 10
0 -2 0 -3
-4
,0
00
REFER TO
FIG. 4-5A
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(5_OF_8).EMF
Figure 4-7. Hover ceiling OGE (sheet 5 of 8)
4-26
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
0
,0
0
12
00
,0
10
00
80
AT
-
AT
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
°C
5
MINIMUM OAT
O
0 10 0
0 -2 0 -3
-4
UM O
MAXIM
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(6_OF_8).WMF
Figure 4-7. Hover ceiling OGE (sheet 6 of 8)
17 DEC 2002
4-27
BHT-407-FM-1
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
14
,0
00
REFER TO
FIG. 4-5A
12
00
80
O
AT
AT
0 -10
MINIMUM OAT
-2
UM O
10
,0
00
MAXIM
0
0 -3
-4
,0
00
14,000 FT HD
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
EL
20
00
20
40
00
10
60
°C
00
-
0
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(7_OF_8).EMF
Figure 4-7. Hover ceiling OGE (sheet 7 of 8)
4-28
17 DEC 2002
TC APPROVED
BHT-407-FM-1
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
00
,0
14
12
,0
00
14,000 FT HD
0
-3
00
,0
10
0
-4
00
80
O
AT
-
60
00
AT
0 0
-2 -1
UM O
°C
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
REFER TO
FIG. 4-5
EL
20
00
40
5
00
0
-40
EXTERNAL = 6000 LB
16
REFER TO
FIG. 4-5A
MAXIM
MINIMUM OAT
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
,0
18
,0
00
20
,0
00
20,000 FT HD
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM-1__FIG_4-7_(8_OF_8).EMF
Figure 4-7. Hover ceiling OGE (sheet 8 of 8)
17 DEC 2002
4-29
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
15,000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
-10
0
OA
T
X
MA
16000
T
OA
14000
10
- °C
5
20
IT
LIM
4
12000
30
10000
3
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
18000
PRESSURE ALTITUDE - FEET
6
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(1_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 1 of 10)
4-30
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
BASIC INLET
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
11,000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
1
2
3
4
5
6
7
8
9
10
20000
MA
XO
LIM
AT
18000
O
AT
-1
-°C
6
0
0
IT
16000
PRESSURE ALTITUDE - FEET
-2
0
-3
0
5
10
14000
4
20
12000
10000
3
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
0
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(2_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 2 of 10)
17 DEC 2002
4-31
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
11,500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
1
2
3
4
5
6
7
8
9
10
-30
-4
0
-2
0
-1
0
18000
OA
T-
16000
0
6
5
°C
14000
MA
AT
12000
4
20
XO
LIM
10000
IT
PRESSURE ALTITUDE - FEET
10
3
30
8000
2
6000
40
PRESSURE ALTITUDE - 1000 METERS
0
20000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(3_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 3 of 10)
4-32
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
BASIC INLET
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
7500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
1
2
3
4
5
6
7
8
9
20000
6
MA
-3
X
OA
LI
T
MI
PRESSURE ALTITUDE - FEET
16000
-2
0
OA
T°C
T
18000
10
-4
0
0
-1
0
5
0
14000
10
4
12000
20
10000
3
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
0
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(4_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 4 of 10)
17 DEC 2002
4-33
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
REDUCE RATE OF CLIMB 170 FT/MIN ABOVE
9000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4000 lb (1814 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
-2
0
-4
16000
-1
OA
14000
T°C
0
5
0
4
10
X
OA
TL
3
20
IT
IM
10000
-3
0
MA
12000
0
8000
30
2
6000
4000
40
PRESSURE ALTITUDE - 1000 METERS
18000
PRESSURE ALTITUDE - FEET
10
1
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(5_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 5 of 10)
4-34
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
BASIC INLET
REDUCE RATE OF CLIMB 170 FT/MIN ABOVE
4500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4000 lb (1814 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
18000
6
A
X
O
A
T
LI
M
PRESSURE ALTITUDE - FEET
16000
OA
T
IT
-4
-°C
-2
0
14000
0
5
-3
0
-1
0
4
0
12000
10
10000
3
20
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
M
10
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(6_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 6 of 10)
17 DEC 2002
4-35
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
REDUCE RATE OF CLIMB 150 FT/MIN ABOVE
6500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
20000
6
7
8
9
10
-4 0
6
18000
-2 0
-1
14000
OA
12000
0
-3
T°C
0
0
4
10
10000
3
AT
XO
8000
20
LIM
IT
6000
4000
30
2
40
1
PRESSURE ALTITUDE - 1000 METERS
5
MA
PRESSURE ALTITUDE - FEET
16000
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(7_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 7 of 10)
4-36
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
BASIC INLET
REDUCE RATE OF CLIMB 150 FT/MIN ABOVE
2000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
OA
T
14000
-4
0
-2
0
-°C
-3
0
4
-1
0
12000
0
10000
3
10
20
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(8_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 8 of 10)
17 DEC 2002
4-37
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
5000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
18000
-4 0
-20
14000
OA
12000
4
-1 0
T°C
-3
0
0
10000
3
10
X
MA
8000
OA
20
2
IT
IM
TL
6000
30
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
10
4000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(9_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 9 of 10)
4-38
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
1000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
-3
0
-4
0
14000
OA
12000
T-
4
-2
0
-1
0
°C
0
10000
3
10
8000
20
2
6000
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-8_(10_OF_10).WMF
Figure 4-8. Rate of climb – takeoff power (sheet 10 of 10)
17 DEC 2002
4-39
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
10,000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
-3
0
-2
18000
0
-1
0
0
O
14000
AT
- °C
10
4
12000
20
3
MA
10000
XO
AT
8000
LIM
30
2
IT
6000
40
4000
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
6
1
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(1_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 1 of 10)
4-40
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
6000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
1
2
3
4
5
6
7
8
9
6
MA
-4
0
XO
AT
18000
-3
0
LIM
-2
IT
16000
PRESSURE ALTITUDE - FEET
10
O
AT
-°
14000
-1
C
5
0
0
4
0
12000
10
10000
3
8000
20
2
6000
PRESSURE ALTITUDE - 1000 METERS
0
20000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(2_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 2 of 10)
17 DEC 2002
4-41
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
7500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
-4
-3
0
0
-2
0
-1
14000
O
12000
5
0
0
AT
-°
C
4
10
10000
3
MA
XO
8000
20
AT
IT
LIM
PRESSURE ALTITUDE - FEET
16000
6000
2
30
PRESSURE ALTITUDE - 1000 METERS
18000
4000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(3_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 3 of 10)
4-42
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
2500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
0
2
3
4
5
6
7
8
9
10
6
XO
AT
IT
LIM
-4
0
16000
-3
0
5
-2
0
O
14000
A
T-
-1
0
°C
4
12000
0
10000
3
10
8000
20
2
6000
PRESSURE ALTITUDE - 1000 METERS
18000
PRESSURE ALTITUDE - FEET
1
MA
20000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(4_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 4 of 10)
17 DEC 2002
4-43
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 170 FT/MIN ABOVE
5000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4000 lb (1814 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
-4
0
-3
0
5
-2
0
-1
14000
O
AT
-
°C
0
4
0
12000
10
10000
3
X
MA
20
IM
TL
OA
8000
2
IT
PRESSURE ALTITUDE - FEET
16000
6000
30
PRESSURE ALTITUDE - 1000 METERS
18000
4000
40
1
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(5_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 5 of 10)
4-44
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 170 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4000 lb (1814 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
-4
14000
O
AT
12000
0
-3
0
-°
C
-1
0
4
-2
0
0
10000
3
10
8000
20
2
6000
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(6_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 6 of 10)
17 DEC 2002
4-45
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 150 FT/MIN ABOVE
3000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
-3
0
-4
0
5
14000
-1
12000
OA
10000
0
4
0
T°C
10
MA
3
XO
AT
8000
20
LIM
IT
PRESSURE ALTITUDE - FEET
-2
0
2
6000
30
PRESSURE ALTITUDE - 1000 METERS
16000
4000
40
-40
1
-20
0
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(7_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 7 of 10)
4-46
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 150 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
4
-4
0
12000
O
AT
-3
0
-°
C
-2
10000
-1
0
3
0
0
8000
10
6000
2
20
1
1800
0
2000
-20
4000
-40
PRESSURE ALTITUDE - FEET
14000
PRESSURE ALTITUDE - 1000 METERS
5
16000
2000
0
0
200
400
600
800
1000
1200
1400
1600
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(8_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 8 of 10)
17 DEC 2002
4-47
BHT-407-FM-1
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
3000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
14000
0
-2
0
4
-1
0
12000
0
OA
T°C
10000
3
10
8000
MA
X
OA
20
TL
PRESSURE ALTITUDE - FEET
-4
-3
0
6000
2
IT
IM
30
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(9_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 9 of 10)
4-48
17 DEC 2002
TC APPROVED
BHT-407-FM-1
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 135 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
-4
12000
-3
0
O
10000
-2
AT
- °C
4
0
0
-1
0
3
0
8000
10
6000
2
20
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FM-1__FIG_4-9_(10_OF_10).WMF
Figure 4-9. Rate of climb – maximum continuous power (sheet 10 of 10)
17 DEC 2002
4-49
BHT-407-FM-1
TC APPROVED
AUTOROTATION GLIDE DISTANCE
20000
18000
Minimum Rate of
Descent Airspeed 55 KIAS
16000
Maximum Glide Distance
Airspeed 80 KIAS
Pressure Altitude (ft)
14000
12000
10000
8000
NOTE: AUTOROTATIONAL DESCENT
PERFORMANCE IS A FUNCTION OF
AIRSPEED AND IS ESSENTIALLY
UNAFFECTED BY DENSITY ALTITUDE
AND GROSS WEIGHT.
6000
4000
2000
0
0
1
2
3
4
5
6
7
8
9
10
11
12
13
14
Distance Over Ground (Nautical Miles)
M407_FM-1__FIG_4-10.EMF
Figure 4-10. Autorotation glide distance
4-50
17 DEC 2002
TC APPROVED
BHT-407-FM-1
AIRSPEED INSTALLATION CORRECTION TABLE
KCAS = (KIAS – INSTRUMENT ERROR – POSITION ERROR)
NOTE: This chart assumes zero instrument error.
KIAS
CLIMB
KCAS
LEVEL FLIGHT
KCAS
20
––
22
30
30
33
40
37
43
50
47
52
60
58
63
70
69
73
80
78
82
90
87
92
100
95
100
110
––
110
120
––
121
130
––
131
140
––
144
M407_FM-1_FIG_4-11.WMF
Figure 4-11. Airspeed installation correction
17 DEC 2002
4-51/4-52
BHT-407-FM-1
Section 5
WEIGHT AND BALANCE
5x
TABLE OF CONTENTS
Subject
Paragraph
INTRODUCTION .........................................................................
EMPTY WEIGHT CENTER OF GRAVITY ..................................
EMPTY WEIGHT .................................................................
CENTER OF GRAVITY .......................................................
GROSS WEIGHT CENTER OF GRAVITY..................................
USEFUL LOADS .................................................................
CENTER OF GRAVITY .......................................................
DOORS OPEN OR REMOVED ...................................................
DOOR WEIGHTS AND MOMENTS ....................................
BALLAST ADJUSTMENT...................................................
COCKPIT AND CABIN LOADING ..............................................
LONGITUDINAL LOADING ................................................
MOST FORWARD AND MOST AFT CG ............................
ALTERNATE LOADING ......................................................
CABIN FLOOR LOADING...................................................
BAGGAGE COMPARTMENT LOADING ...................................
FUEL LOADING ..........................................................................
SAMPLE LOADING PROBLEM .................................................
5-1....................
5-2....................
5-2-A................
5-2-B................
5-3....................
5-3-A................
5-3-B................
5-4....................
5-4-A................
5-4-B................
5-5....................
5-5-A................
5-5-B................
5-5-C................
5-5-D................
5-6....................
5-7....................
5-8....................
Page
Number
5-3
5-3
5-3
5-3
5-3
5-3
5-3
5-4
5-4
5-4
5-4
5-5
5-5
5-5
5-5
5-5
5-6
5-6
LIST OF FIGURES
Title
Figure
Number
Fuselage stations.......................................................................
Buttock lines ..............................................................................
Fuel center of gravity.................................................................
5-1....................
5-2....................
5-3....................
Page
Number
5-7
5-8
5-9
LIST OF TABLES
Title
Table
Number
Door Weights and Moments (U.S.) ...........................................
Cabin and baggage loading (U.S.)............................................
Cabin and baggage loading (Metric) ........................................
Fuel Loading (U.S.) ....................................................................
Fuel Loading (Metric).................................................................
5-1...................
5-2...................
5-3...................
5-4...................
5-5...................
Page
Number
5-10
5-11
5-12
5-13
5-14
17 DEC 2002
5-1
BHT-407-FM-1
LIST OF TABLES (CONT)
Title
Table
Number
Fuel Density vs Temperature....................................................
Sample Loading Problem (U.S.) ...............................................
Sample Loading Problem (Metric)............................................
Weight and Balance Worksheet (U.S.).....................................
Weight and Balance Worksheet (Metric) .................................
5-6 ..................
5-7 ..................
5-8 ..................
5-9 ..................
5-10 ................
5-2
17 DEC 2002
Page
Number
5-15
5-16
5-17
5-18
5-19
BHT-407-FM-1
Section 5
WEIGHT AND BALANCE
5
5-1.
INTRODUCTION
This section presents loading information and
instructions necessary to ensure that flight
can be performed within approved gross
weight and center of gravity limitations as
defined in Section 1.
5-2.
5-2-A.
EMPTY WEIGHT CENTER OF
GRAVITY
EMPTY WEIGHT
The empty weight condition 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
em pty we ight and ce nter of g ravity are
recorded on the Actual Weight Record, a copy
of which should be carried in the helicopter to
enable weight and balance computations.
5-2-B.
CENTER OF GRAVITY
An empty weight center of gravity chart is
provided in maintenance manual as a guide to
simplify computing ballast requirements. This
chart was derived from gross weight
longitudinal center of gravity limits shown in
Section 1, using most forward and most aft
useful loads for standard seating and fuel.
NOTE
Empty weight center of gravity chart
is not valid if helicopter has a
nonstandard fuel system or seating
arrangement.
5-3.
GROSS WEIGHT CENTER OF
GRAVITY
Gross weight condition is empty weight
condition plus useful load.
5-3-A.
USEFUL LOADS
Useful load consists of usable fuel, engine oil,
c re w, pa s s e ng e r s, ba g g a g e a n d ca r g o.
Combinations of these items which have most
adverse effect on helicopter center of gravity
are known as most forward and most aft
useful loads. Whenever cargo and/or baggage
a re ca rr ie d , th es e us e fu l lo a ds m a y b e
different for each flight, and weight and
balance must be computed to ensure gross
weight and center of gravity will remain within
limits throughout flight.
Standard most forward and most aft useful
loads are combinations of fuel, crew and
passenger loading only. These loads, in
conjunction with empty weight center of
gravity chart, allow passengers only (no
baggage or other cargo) to be carried within
a p p r o p ri a t e w e i g h t l im i ta t io n s w it h o u t
computing center of gravity for each flight.
If helicopter has a nonstandard fuel system or
seating arrangement, or is not ballasted in
accordance with empty weight center of
gravity chart in maintenance manual, pilot
must determine weight and balance to ensure
gross weight and center of gravity will remain
within limits throughout each flight.
5-3-B.
CENTER OF GRAVITY
It is the responsibility of the pilot to ensure
that helicopter is properly loaded to maintain
17 DEC 2002
5-3
BHT-407-FM-1
center of gravity throughout each flight within
gross weight center of gravity limits shown in
Section 1 or appropriate supplement. Gross
weight longitudinal and lateral center of
gravity can be calculated using Actual Weight
Record, diagrams and loading tables in this
section and loading tables in applicable flight
manual supplements.
Example:
When removing a left door only, subtract
positive weight value and negative moment
value shown in table. Net effect on helicopter
is a reduction in weight and a shift in lateral
CG to right (positive direction).
5-4-B.
When carrying baggage, cargo or
n o n s t a n d a r d l o a d s , e f f e c ts o f f u e l
consumption and addition/deletion of
pass eng ers , b ag gag e or c arg o at
intermediate points should be checked prior
to flight.
Following check can be made to determine if
a ballast adjustment is necessary after doors
are removed or installed.
1.
For helicopters without ballast or
with nose ballast, apply weight and
moment changes to most aft useful
load condition to determine if an
increase in nose ballast is required,
or a reduction is allowed.
2.
For helicopters with tail ballast,
apply weight and moment changes
to most forward useful load
condition to determine if a reduction
i n ta i l b a l la s t i s a ll o w e d , o r a n
increase is required.
Significant fuselage stations and buttock
lines are shown in Figures 5-1 and 5-2 to aid
in weight and balance computations.
5-4.
DOORS OPEN OR REMOVED
When one or more cabin doors are removed,
helicopter may exceed gross weight center of
gravity limits during flight. If using Weight
empty center of gravity chart, refer to BHT407-MM-1, a ballast adjustment to offset
moment change is necessary (Table 5-1).
Otherwise, gross weight center of gravity
should be computed for each flight.
5-4-A.
DOOR WEIGHTS AND MOMENTS
Following table presents weight and moment
adjustments for cabin doors. Sign convention
for buttock lines used to compute lateral
moments are:
1.
Left is negative.
2.
Right is positive.
ACTION
Remove
Install
5-4
MOMENT CHANGE
LEFT DOOR
RIGHT DOOR
Positive (+)
Negative (-)
Negative (-)
Positive (+)
17 DEC 2002
BALLAST ADJUSTMENT
NOTE
Ballast changes are performed by
maintenance personnel. After any
ballast change, Actual Weight Record
must be revised to show new weight
empty condition.
5-5.
COCKPIT AND CABIN
LOADING
Loading tables (Tables 5-2 and 5-3) provide
weights and moments for each passenger
loc ation , litte r patie nt a nd bag ga ge
compartment in both U.S. and metric units.
To find moments for weights in excess of
those shown on tables, multiply weight by
fuselage station at which center of gravity of
BHT-407-FM-1
the object is located. An alternate method is
to calculate amount of weight in excess of
maximum weight listed on table, then read
moment for this excess weight from table and
add it to moment for maximum weight shown
on table. This will give desired moment for the
object.
5-5-A.
1.
LONGITUDINAL LOADING
A minimum weight of 170 pounds
(77.1 kilograms) is required in
cockpit at fuselage station 65.0
when the empty weight center of
gravity chart is used.
5-5-C.
ALTERNATE LOADING
Gross weight center of gravity chart must be
used to determine cabin loading requirements
under following conditions:
1.
Whenever cargo and/or baggage are
carried.
2.
When actual passenger weights are
used.
3.
When seating arrangement and/or
fuel system are non-standard.
4.
When performing specialty missions,
such as hoisting or rappelling.
2.
Passenger seating is unrestricted.
5-5-D.
3.
Cargo loading is restricted only by
floor load limit. Refer to Section 1.
Cabin floor is structurally designed for 75
pounds per square foot (3.7 kilograms per 100
square centimeters).
5-5-B.
MOST FORWARD AND MOST
AFT CG
When using empty weight center of gravity
chart, following combinations of crew, fuel
a n d pa s s e n g e r l o a d i n g w i l l h a v e m o s t
extreme effects on longitudinal center of
gravity, assuming standard weights for all
crew and passengers.
1.
Most forward CG will occur with
forward and mid seats occupied and
fuel quantity of 74.8 gallons (283.0
liters).
2.
Most aft CG will occur with one
forward seat occupied (pilot) and
fuel quantity of 28.4 gallons (107.5
liters).
Since center of gravity of aft passengers is on
aft limit, weight of passengers is not included
in most aft useful load. However when most
aft center of gravity of a configuration is
forward of aft limit, addition of aft passengers
will shift center of gravity further aft, and
should be included in computation.
5-6.
CABIN FLOOR LOADING
BAGGAGE COMPARTMENT
LOADING
When weight is loaded into baggage
compartment, the pilot is required to compute
weight and balance, regardless of passenger
loading.
B a g g a g e c o m pa r t m e n t i s s t r u c t u r a l l y
designed for 86 pounds per square foot (4.2
kilograms per 100 square centimeters) for a
total weight of 250 pounds (113.4 kilograms).
Loading of baggage compartment should be
from front to rear. Load shall be secured to
tiedown fittings if shifting of load in flight
could result in structural damage to baggage
compartment or in gross weight center of
gravity being exceeded.
If load is not secured, center of gravity must
be computed with load in most adverse
position.
17 DEC 2002
5-5
BHT-407-FM-1
5-7.
FUEL LOADING
Longitudinal center of gravity of fuel shifts as
it is consumed (Figure 5-3). Extreme effects of
fuel consumption on helicopter center of
gravity fo r stand ard fuel system are as
follows:
1.
Critical fuel for computing most
forward useful load is 74.8 gallons
(283.0 liters).
2.
Critical fuel for computing most aft
useful load is 28.4 gallons (107.5
liters).
Fuel loading tables (Tables 5-4 and 5-5) list
usable fuel quantities, weight and moments in
both U.S. and metric units.
Fuel density vs temperature (Table 5-6), is
provided to calculate fuel weight variation for
equivalent volumes of fuel caused by a
change in temperature. For example weight of
5-6
17 DEC 2002
127.8 gallons (full fuel) of JP-5 at -40°F is
913.8 pounds (414.5 kilograms) versus 869.0
pounds (394.1 kilograms) shown on Fuel
loading chart (Tables 5-4 and 5-5).
5-8.
SAMPLE LOADING
PROBLEM
A sample loading problem showing derivation
of critical gross weights and center of gravity
locations for a typical mission is presented in
U.S. and metric units (Tables 5-7 and 5-8).
Method shown derives a gross weight with
ze ro f ue l f or e ac h lo ad co nd itio n to b e
checked, then adds appropriate fuel weight
and moment read directly from Fuel loading
table. Center of gravity for each condition is
calculated by dividing total moment by total
weight.
Forms have been provided (Tables 5-9 and 510) in both U.S. and Metric Units, to aid in
computing critical load conditions for a flight.
BHT-407-FM-1
Figure 5-1. Fuselage stations
17 DEC 2002
5-7
BHT-407-FM-1
Figure 5-2. Buttock lines
5-8
17 DEC 2002
BHT-407-FM-1
Figure 5-3. Fuel center of gravity
17 DEC 2002
5-9
BHT-407-FM-1
Table 5-1. Door Weights and Moments (U.S.)
5-10
17 DEC 2002
BHT-407-FM-1
Table 5-2. Cabin and baggage loading (U.S.)
17 DEC 2002
5-11
BHT-407-FM-1
Table 5-3. Cabin and baggage loading (Metric)
5-12
17 DEC 2002
BHT-407-FM-1
Table 5-4. Fuel Loading (U.S.)
17 DEC 2002
5-13
BHT-407-FM-1
Table 5-5. Fuel Loading (Metric)
5-14
17 DEC 2002
BHT-407-FM-1
Table 5-6. Fuel Density vs Temperature
17 DEC 2002
5-15
BHT-407-FM-1
Table 5-7. Sample Loading Problem (U.S.)
5-16
17 DEC 2002
BHT-407-FM-1
Table 5-8. Sample Loading Problem (Metric)
17 DEC 2002
5-17
BHT-407-FM-1
Table 5-9. Weight and Balance Worksheet (U.S.)
5-18
17 DEC 2002
BHT-407-FM-1
Table 5-10. Weight and Balance Worksheet (Metric)
17 DEC 2002
5-19/5-20
BHT-407-FM-1
Appendix A
OPTIONAL EQUIPMENT SUPPLEMENTS
A
TABLE OF CONTENTS
Subject
Paragraph
Number
Page
Number
Optional Equipment...............................................................................
A-1 ..........
A-3
Table
Number
Page
Number
LIST OF TABLES
Subject
Flight Manual Supplements for Optional Equipment .........................
A-1 .......
A-3
19 FEB 2007—Rev. 5———A-1/A-2
BHT-407-FM-1
Appendix A
OPTIONAL EQUIPMENT SUPPLEMENTS
A
A-1. OPTIONAL EQUIPMENT
changes, additions, improvement, etc., on its
products previously manufactured.
Bell Helicopter Textron's policy is one of
continuous product improvement and Bell
reserves the right to incorporate changes,
make additions to and improve its products
without imposing any obligation upon the
c o m pa n y t o fu r n i s h f o r o r i n s ta l l s u c h
The following items may be installed on the
basic helicopter by authorized personnel.
Only the optional equipment listed in
Table A-1 require a Flight Manual Supplement.
Table A-1: Flight Manual Supplements for Optional Equipment
CURRENT
REVISION
NAME OF EQUIPMENT
KIT NUMBER
DATE
CERTIFIED
BHT-407-FMS-1
Lightweight Emergency
Flotation Landing Gear
407-706-008
11 APR 96
Original
BHT-407-FMS-2
High Skid Gear
407-706-007
14 FEB 96
Original
BHT-407-FMS-3
Particle Separator
206-706-212
1 MAR 96
Reissue
16 DEC 02
BHT-407-FMS-4
Snow Deflector
206-706-208
1 MAR 96
Reissue
16 DEC 02
BHT-407-FMS-5
Cargo Hook
206-706-341
14 FEB 96
Rev. 1
4 SEP 98
BHT-407-FMS-6
Auxiliary Fuel Kit
407-706-011
20 MAR 96
Original
BHT-407-FMS-7
Litter(s)
407-706-631
or
407-799-100
or
407-799-001
14 FEB 96
Rev. 1
16 SEP 99
BHT-407-FMS-8
Reserved
BHT-407-FMS-9
Reserved
BHT-407-FMS-10
Helicopters Registered in U.S.A.
Canceled
19 FEB 2007—Rev. 5———A-3
BHT-407-FM-1
Table A-1: Flight Manual Supplements for Optional Equipment
NAME OF EQUIPMENT
KIT NUMBER
BHT-407-FMS-11
Reserved
BHT-407-FMS-12
Reserved
BHT-407-FMS-13
Reserved
BHT-407-FMS-14
Reserved
BHT-407-FMS-15
Reserved
BHT-407-FMS-16
Reserved
BHT-407-FMS-17
Cargo Tie-down Provisions Kit
407-705-201
BHT-407-FMS-18
Reserved
BHT-407-FMS-19
Reserved
DATE
CERTIFIED
CURRENT
REVISION
1 APR 96
Original
BHT-407-FMS-20
KLN 89B GPS Navigator
407-705-001
14 FEB 96
Rev. 1
26 NOV 96
BHT-407-FMS-21
Fire Detection System
407-799-004
or
407-706-015
or
407-706-025
2 MAY 96
Reissue
7 JUL 04
BHT-407-FMS-22
Auxiliary Vertical Fin Strobe
Lights
407-899-023
10 MAY 96
Original
BHT-407-FMS-23
Ryan Traffic Collision Avoidance
Device
407-899-022
15 MAY 96
Original
8 MAY 98
Reissue
17 DEC 02
BHT-407-FMS-24
BHT-407-FMS-25
Quiet Cruise Mode
Reserved
407-706-016
BHT-407-FMS-26
Modified Hydromechanical Unit
Canceled
BHT-407-FMS-27
FADEC Software Version 5.201
Canceled
BHT-407-FMS-28
Increased Internal Gross Weight
407-706-020
BHT-407-FMS-29
Airspeed Actuated Pedal Stop
BHT-407-FMS-30
A-4———Rev. 5—19 FEB 2007
16 MAR 99
Reissue
16 DEC 02
Canceled
Reserved
BHT-407-FM-1
Table A-1: Flight Manual Supplements for Optional Equipment
NAME OF EQUIPMENT
BHT-407-FMS-31
Increased APC Starter
Generator Load
BHT-407-FMS-32
Hanger Bearing Vibration
Monitor Kit
(not distributed to all
customers)
DATE
CERTIFIED
407-706-026
15 JUN 05
Original
407-706-023
15 JUL 03
Original
08 JAN 02
Original
20 MAY 99
Original
BHT-407-FMS-CAA
United Kingdom
Registered Helicopters
BHT-407-FMS-IAC AR
Interstate Aviation Committee
— Aviation Register
Commonwealth of Independent
States
CURRENT
REVISION
KIT NUMBER
407-706-021
19 FEB 2007—Rev. 5———A-5/A-6
BHT-407-FM-I-ATO
11MODEL 407
ROTORCRAFT
FLIGHT MANUAL
FOR REPUBLIC OF PHILIPPINES
REGISTERED HELICOPTERS
Republic of the Philippines
Department of Transportation and Communications
AIR TRANSPORTATION OFFICE
NAIA, 1300 Pasay City, M.M
VALIDATED AIRCRAFT TYPE CERTIFICATE
NO. H1TSN-001
Inspect And Evaluated By:
L. t&JERA, JR.
in er III
Reviewed By:
JO9E'C. ROBLES, JR.
Chief, Aircraft Engineering Section
(RET.)
6ATURNINO B. DELA
Chief, Aviation Safety
DATE: November 12, 1996
THE AVLAT70N REGULATORY AUTHORITY FOR THIS FLIGHT MANUAL IS THE
AIR TRANSPORTATION OFFICE OF THE RUPUBLIC OF THE PHILIPPINES.
R.P. REGISTERED HELICOPTERS ARE APPROVED BY ATO IN ACCORDANCE
WITH THE PROVISIONS OF REPUBLIC ACT 776 SERIES OF 1952 AS AMENDED BY
PRESIDENTIAL DECREE. NO. 844. 1278. 1462 AND EXECUTIVE ORDER NO. 546.
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
Bell Helicopter
A Subsidiary of Textron Inc
COPYRIGHT NOTICE
COPYRIGHT
1996
BELL @ HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON,
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
POST OFFICE BOX 482
FORT WORTH, TEXAS 76101
12 NOVEMBER 1996
BHT 407-FM-I-ATO
NOTICE PAGE
The pages contained herein shall be attached to the basic flight Manual
when the helicopter is registered in the Republic of the Philippines.
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
BHT-407-FM-I-ATO
LOG OF REVISIONS
Original
..........
0
...
November 12, 1996
LOG OF PAGES
PAGE
REVISION
NO.
PAGE
REVISION
NO.
Title - NP ............................ 0
A/B ................................... 0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages: dispose of
superseded pages
A/B
BHT-407-FM-CTA
9mro
MODEL
ROTORCRAFT
FLIGHT MANUAL
4, -' -
REGISTRATION NO.
0
m
DOCUMENT NO.
SERIAL NO.
FOREIGN AUNWR" REPRESENTATIVE
DIRECTOR
AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
THIS MANOL AND ASSOCIATED SUPPLEMENTS ARE APPROVED
IN
ACCORDANCE WITH SECTION 21.29 OF RBHA 21 FOR BRAZILIAN REGISTERED
AIRCRAFT AND IS APPROVED BY THE D.O.T. ON BEHALF OF THE CENTRO
TECNICO AEROESPACIAL, WHEN HELICOPTER IS REGISTERED IN BRAZIL.
Tom
THIS AIRCRAFT SHOULD BE OPERATED IN ACCORDANCE WITH THE
LIMITATIONS AND INSTRUCTIONS HEREIN ESTABLISHED.
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
Bell Helicopter
COPYRIGHT NOTICE
COPYRIGHT
1998
BELL 0 HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
A Subsidiary of Textron Inc.
POST OFFICE BOX 482
FORT WORTH, TEXAS 76101
CTA APPROVED - 3 APRIL 1998
BHT 407-FM-CTA
NOTICE PAGE
°._
.be
The following Warning is not applicable to helicopters on which all kits and
customizing installations have been qualified and approved by Bell Helicopter.
WARNING
=o°D
H}m
THE HELICOPTER MAY CONTAIN INSTALLATIONS, PARTS, OR
==!
PROCESSES CERTIFIED BY PARTIES OTHER THAN BELL
--1
HELICOPTER TEXTRON. BELL HELICOPTER CAN NOT
D=m
v-1
CONFIRM THAT SUCH INSTALLATIONS HAVE BEEN FULLY
QUALIFIED OR CONFORMED TO BELL HELICOPTER DESIGN
CRITERIA. AS A RESULT OF SUCH INSTALLATIONS, BELL
HELICOPTER SUPPLIED DATA MAY NOT BE VALID
i-1
CONCERNING IN-FLIGHT HANDLING QUALITIES, WEIGHT AND
Nam
<_O
BALANCE, OR HELICOPTER PERFORMANCE. IF MULTIPLE
STC KITS OR SIMILAR INSTALLATIONS ARE INCORPORATED,
.Or
F-}
THERE MAY BE NO VALID TEST DATA TO QUALIFY THE
HELICOPTER AS MODIFIED BY THESE INSTALLATIONS. FOR
REVISED DATA, CONTACT THE OWNER OF THE INSTALLED
STC OR THE SUPPLIER FOR THE APPLICABLE APPROVAL OF
EACH INSTALLATION.
O':.
___
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
r-1r
MODEL
BHT-407-FM-IAC AR
ROTORCRAFT
FLIGHT MANUAL
DOCUMENT N0.
REGISTRATION N0.
SERIAL NO.
DATE
DIRECTOR - AI
DEPARTM
L1
.
lf
AFT CERTIFICATION BRANCH
T OF TRANSPORT, CANADA
THIS MANUAL AND ASSOCIATED SUPPLEMENTS ARE APPROVED !N
ACCORDANCE WITH SECTIONS 2.10 AND 4.7 OF AP 21 FOR C1S REGISTERED
AIRCRAFT AND IS APPROVED BY THE D.O.T. ON BEHALF OF THE AVIATION
REGISTER OF THE INTERSTATE AVIATION COMMITTEE.
THIS AIRCRAFT SHOULD BE OPERATED !N ACCORDANCE WITH THE
Z
LIMITATIONS AND INSTRUCTIONS HEREIN ESTABLISHED.
COPYRIGHT NOTICE
DC)
z
-0c
_l-_>
r-0
c)
o)
F-J=
1999
COPYRIGHT
BELL 0 HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell Helicopter
A Subsidiary of Textron Inc.
o
m
z
o
-I
)
THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS
POST OFFICE 80X 482
FORT WORTH, TEXAS 76101
20 MAY 1999
BHT-407-FM-IAC AR
NOTICE PAGE
The following Warning is not applicable to helicopters on which all kits and
customizing installations have been qualified and approved by Bell Helicopter.
WARNING
THE HELICOPTER MAY CONTAIN INSTALLATIONS, PARTS,
OR PROCESSES CERTIFIED BY PARTIES OTHER THAN
BELL HELICOPTER TEXTRON. BELL HELICOPTER CAN
NOT CONFIRM THAT SUCH ITEMS HAVE BEEN FULLY
QUALIFIED OR CONFORMED TO BELL HELICOPTER
DESIGN CRITERIA. BELL HELICOPTER SUPPLIED DATA
MAY NOT BE VALID CONCERNING IN-FLIGHT HANDLING
QUALITIES, WEIGHT AND BALANCE, OR SIMILAR
D=T
INSTALLATIONS. THERE MAY BE NO VALID TEST DATA TO
QUALIFY THE HELICOPTER AS MODIFIED BY THESE ITEMS.
z0--i
FOR REVISED DATA, CONTACT THE OWNER OF THE
INSTALLED STC (OR EQUIVALENT CERTIFICATION) OR THE
SUPPLIER FOR THE APPLICABLE APPROVAL OF EACH
INSTALLATION.
PROPRIETARY RIGHTS NOTICE
Manufacturer's Data portion of this manual is proprietary to Bell
W-0
(CND
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.
®000)
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
B HT-407-FM S-i
'!407
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
LiGHTWEIGHT EMERGENCY
FLOTATION LANDING GEAR
407-706-008
CERTIFIED
11 APRIL 1996
This supplement shall be attached to Model 407
Flight Manual when 407-706-008 Lightweight
Emergency Flotation Landing Gear 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
1996
COPYRIGHT
BELL HELICOPTER INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell Helicopterti4it.1i
A Subsidiary of Textron Inc.
ST OFFICE BOX 482 • FORT WORTH. TEXAS 78101
11 APRIL 1 99
BHT-407-FMS-1
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-407-FMS-1
LOG OF REVISIONS
Original ........... 0
11 Apr 1996
LOG OF PAGES
REVISION
NO.
PAGE
FLIGHT MANUAL
Title—NP
A—B
PAGE
C/D
0
0
i/li
1—4
5/6
REVISION
NO.
0
0
0
0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
A
BHT-407-FMS-1
DOT APPROVED
LOG OF APPROVED REVISIONS
Original
.0
.11 Apr 1996
APPROVED:"'I0,
PCHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
B
DATE: /1 ,'9pe q"
BHT-407-FMS-1
LOG OF FAA APPROVED REVISIONS
Original
.0
.11 Apr 1996
C/D
BHT-407-FMS-1
GENERAL INFORMATION
Lightweight emergency flotation landing
gear kit (407-706-008) will allow helicopter
to land in water during an emergency
situation. Kit consists of six skid mounted
pop-out float bags, an inflation system
with electrically operated solenoid valves,
and attaching hardware. Two pneumatic
charging bottles, located on underside of
helicopter, are interconnected by a
pneumatic line that will cause charging
bottle valve to open in event that its
solenoid valve fails while floats are being
inflated. Each float assembly is equipped
with an inlet check valve, high pressure
relief valve which opens at 5.25 PSIG
0.25 PSI and a finger operated manual
stop-cocklinflation valve. A GEN FAIL
caution light alerts pilot of generator
failure and of battery power possibly being
insufficient to inflate floats. Float inflation
time is approximately 5 seconds.
i/u
DOT APPROVED
Lsection
BHT-407-FMS-1
1 1
LIMITA TIONS
1-3. TYPES OF OPERATION
Emergency floats are installed for
Maximum allowable airspeed, floats
inflated, is 60 KIAS.
Maximum autorotation airspeed, floats
inflated, is 60 KIAS.
assistance during emergency ditching.
1-8. ALTiTUDE
1-6. WEiGHT AND CENTER OF
GRAVITY
Maximum inflation altitude is 5000 feet H.
1-9. MANEUVERING
Actual weight change shall be determined
after kit is installed and ballast readjusted,
if necessary, to return empty weight CG to
within allowable limits. Refer to Center of
gravity vs weight empty chart in BHT-407MM-i.
1-9-B. CLIMB AND DESCENT
Maximum rate of climb with floats inflated
is 1000 feet per minute.
1-20. INSTRUMENT
1-7. AIRSPEED
1-7-A. FLOATS STOWED
MARK1NGS AND PLACARDS
FLOAT
ARMINGJNFLATLON
ABOVE 60 KIAS
PROHIBITED
Floats stowed, covers installed — Same as
basic helicopter.
1-7-B. FLOATS INFLATED
Maximum arming/inflation airspeed is 60
KIAS.
1
BHT-407-FMS-1
DOT APPROVED
Section 2
NORMAL PROCEDURES
2-3. PREFLIGHT CHECK
1. Floats — Stowed.
2. Nitrogen lines — Condition and
security.
3. Float covers — Clean and
secured.
4. Float inflation cylinders — Check
for proper inflation pressure vs
4. FLOAT TEST and FLOAT ARM
lights — Press C/W LT TEST
button.
5. FLOAT TEST button — Press and
hold.
6. FLOAT INFLATE button — Press;
check FLOAT TEST light
illuminates. Release button,
check light extinguishes.
temperature and altitude. Refer to
7. FLOAT TEST button — Release.
electrical connectors for security.
8. FLOAT ARM switch — Up, guard
open. Check FLOAT ARM light
placard on cylinders. Check
2-4. INTERIOR AND
PRESTART CHECK
2-4-A. PREFLIGHT FLOAT SYSTEM
CHECK
1. BATT switch — BATT. With GEN
switch OFF, verify GEN FAIL light
illuminates.
illuminates, then switch down,
guard closed. Check light
extinguishes.
2-9. IN-FLIGHT OPERATIONS
2-9-A. OVER WATER
OPERATIONS
1. FLOAT ARM switch — Up, guard
CAUTION
;4,444444,4,,4444
open.
2. FLOAT ARM light — illuminated.
IF GEN FAIL LIGHT DOES NOT
ILLUMINATE, MONITOR
VOLTMETER TO DETERMINE
GENERATOR OPERATION. IF
VOLTAGE DROPS BELOW 25
VOLTS, PERFORM GEN FAIL
CORRECTIVE ACTION PER TABLE
3-1.
2. FLOAT ARM switch — Down,
guard closed.
3. FLOATS circuit breaker — Check
in.
2
CAUTION
DURING FLIGHT AT ALTITUDES
ABOVE 500 FEET AGL AND AT
AIRSPEEDS OF 60 KIAS AND
ABOVE, SYSTEM SHOULD BE
DEACTIVATED BY PLACING
FLOAT ARM SWITCH TO DOWN
POSITION AND CLOSING GUARD.
3. Rearm system prior to landing.
BHT-407-FMS-1
DOT APPROVED
2-9-B. OVER LAND OPERATiONS
FLOAT ARM switch — Down, guard
AUTOROTATIONS SHALL BE
AVOIDED DUE TO NOSEDOWN
PITCHING.
closed.
RUN-ON LANDINGS, ON OTHER
THAN A HARD FIRM SURFACE,
SHOULD BE EXERCISED WITH
2-10. DESCENT AND
CAUTION.
LANDING
NOTE
Tail-low run-on landings should
WARNIN
be avoided to prevent nosedown
pitching.
IF CG IS AFT OF STATION 126,
PRACTICE TOUCHDOWN
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3-1. INTRODUCTION
WARNING
Table 3-1 presents fault conditions and
corrective actions for cautions lights.
IF GEN FAIL LIGHT ILLUMINATES,
BATTERY POWER MAY NOT BE
SUFFICIENT TO INFLATE
FLOATS. IF VOLTAGE DROPS
BELOW 25 VOLTS, PERFORM
GEN FAIL CORRECTIVE ACTION
PER TABLE 3-1.
Table 3-1.
PANEL
WORDING
FAULT
CONDITION
CORRECTIVE ACTION
GEN FAIL
Generator not
connected to
DC buss.
Verify fault with AMP or
VOLT gage. Over land: GEN
switch — RESET, then ON. If
GEN FAIL light remains
illuminated or voltage drops
below 25 volts, switch —
OFF. Land as soon as
practical.
3
DOT APPROVED
BHT-407-FMS-1
Table 3-1.
FAULT
CONDITION
PANEL
WORDING
(Cont)
CORRECTIVE ACTION
Over water: GEN switch —
RESET, then ON. If GEN FAIL
light remains illuminated or
voltage drops below 25 volts,
switch — OFF. Turn oft all
nonessential electrical
equipment to conserve
battery power. Land as soon
as practical.
3-15. EMERGENCY FLOAT
4. FLOAT ARM light — illuminated.
INFLATION
CAUTION
1. Reduce airspeed below maximum
inflation airspeed — 60 KIAS.
2. Establish autorotation or low
power descent at approximately
MAXIMUM INFLATION ALTiTUDE
IS 5000 H.
5. FLOAT INFLATE button — Press.
500 feet per minute.
NOTE
If floats are inflated in level flight,
there is a possibility that floats
will not align, which will allow
right or left forward bag to
oscillate. If this occurs, a low
power descent will align float
bags and stop oscillation.
3. FLOAT ARM switch — Up, guard
open.
4
3-16. AFTER EMERGENCY
WATER LANDING
WARN
FLIGHT FOLLOWING A WATER
LANDING IS PROHIBITED.
BHT-407-FMS-1
DOT APPROVED
Section 4
PERFORMA NCE
4-5. HOVER CEILING
45A. INGROUNDEFFECT
HOVER
4-5-B. OUT-OF-GROUND-EFFECT
HOVER
Out-of-ground-effect hover performance is
same as basic helicopter.
Subtract 50 pounds (22.68 kilograms) from
IGE hover gross weight for takeoff power
or maximum continuous power.
5/6
B HT-407-FMS-2
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
HIGH SKID GEAR
407-706-007
CERTIFIED
14 FEBRUARY 1996
This supplement shall be attached to Model 407
Flight Manual when High Skid Gear 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
BELL® HELICOPTER INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISRDN OF TEXTRON CANADA LTD.
Bell Helicopterki *4 ti.11
A Subsidiary 01 Textron Inc.
POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101
14 FEBRUARY 1996
BHT-407-FMS-2
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-407-FMS-2
LOG OF REVISIONS
Original
.0 . 14 Feb 96
LOG OF PAGES
REVISION
NO.
PAGE
C/D
FLIGHT MANUAL
Title—NP
A—B
PAGE
0
0
i/li
1—2
REVISION
NO.
0
0
0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
A
BHT-407-FMS-2
DOT APPROVED
LOG OF APPROVED REVISIONS
Original .......... 0 . 14 Feb 96
APPROVED:
1<
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
B
DATE: /L Fe6 9
BHT-407-FMS-2
LOG OF FAA APPROVED REVISIONS
Original
.0 . 14 Feb 96
C/D
DOT APPROVED
BHT-407-FMS-2
GENERAL INFORMATION
High skid landing gear kit (407-706-007)
provides approximately 8.75 inches (222.25
millimeters) of additional ground clearance
over standard skid gear.
i/il
BHT-407FMS-2
DOT APPROVED
[ Section 1
LIMITA TIONS
1-6. WEIGHT AND CENTER OF
GRAVITY
if necessary, to return empty weight CG to
within allowable limits. Refer to Center of
gravity vs weight empty chart in Bl-IT-407-
MM-i.
Actual weight change shall be determined
after kit is installed and ballast readjusted,
[ Section 2
NORMAL PROCEDURES
2-10. DESCENT AND
WARNING
LANDING
Tail-low run-on landings should be
avoided to prevent nosedown pitching.
RUN-ON LANDINGS ON OTHER
THAN A HARD, FIRM SURFACE
SHOULD BE EXERCISED WITH
CAUTION.
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
I
No change from basic manual.
1
BHT-407-FMS-2
DOT APPROVED
L Section 4
PERFORMA NCE
4-5. HOVER CEILING
4-5-B. OUT- OF-GROUND-EFFECT
HOVER
45A. INGROUNDEFFECT
HOVER
Subtract 50 pounds (22.68 kilograms) from
IGE hover gross weight for takeoff power
or maximum continuous power.
2
Out-of-ground-effect hover performance is
same as basic helicopter.
BHT-407-FMS-3
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
PARTICLE SEPARATOR
206-706-212
CERTIFIED
1 MARCH 1996
This supplement shall be attached to Model 407 Flight
Manual when Particle Separator 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.
COPYRIGHT NOTICE
2002
COPYRIGHT
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD. ALL RIGHTS
RESERVED
POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101
REISSUED — 16 DECEMBER 2002
BHT-407-FMS-3
NOTICE PAGE
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
16 DEC 2002
BHT-407-FMS-3
LOG OF REVISIONS
Original ....................... 0................... 01 MAR 96
Revision ...................... 1................... 01 MAY 97
Revision ...................... 2................... 04 SEP 98
Revision ......................3 ................... 17 APR 00
Reissue........................0 ................... 16 DEC 02
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
MANUFACTURER’S DATA
FLIGHT MANUAL
Title............................................................ 0
NP .............................................................. 0
A — B ........................................................ 0
C/D............................................................. 0
i/ii ............................................................... 0
1 — 38 ....................................................... 0
39/40 .......................................................... 0
REVISION
NO.
41/42 ..........................................................0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
16 DEC 2002
A
BHT-407-FMS-3
LOG OF TC APPROVED REVISIONS
Original ........................0 ................... 01 MAR 96
Revision.......................1 ................... 01 MAY 97
Revision.......................2 ................... 04 SEP 98
Revision.......................3 ................... 17 APR 00
Reissue ........................0 ................... 16 DEC 02
B
16 DEC 2002
BHT-407-FMS-3
LOG OF FAA APPROVED REVISIONS
Original ....................... 0................... 01 MAR 96
Revision ...................... 1....................01 MAY 97
Revision ...................... 2....................18 MAY 99
Revision ...................... 3 ................... 05 MAY 00
Reissue........................ 0 .................... 24 SEP 03
16 DEC 2002
C/D
BHT-407-FMS-3
GENERAL INFORMATION
Bell particle separator kit (206-706-212)
consists of particle separator, bleed air tubing
and hose, electrical cable and required
hardware for installation.
This supplement incorporates performance
information for various combinations of Bell
kits. It also includes limitations and operating
procedures made necessary because of kit
combinations. This supplement is not
intended to replace approved supplements for
other optional equipment, but should be used
in conjunction with such supplements.
16 DEC 2002
i/ii
TC APPROVED
BHT-407-FMS-3
Section 1
LIMITATIONS
1
1-3. TYPES OF OPERATION
Particle separator can be removed and the
engine air intake screen installed to attain
basic helicopter performance.
1-5. CONFIGURATION
1-5-A. OPTIONAL EQUIPMENT
F o r o p e r a t io n s w it h pa r t i c le s e pa r a t o r
installed in conjunction with 206-706-208
snow deflector, refer to LIMITATIONS section
and PERFORMANCE section of snow
deflector supplement (BHT-407-FMS-4).
1-6. WEIGHT AND CENTER OF
GRAVITY
Actual weight change shall be determined
after kit is installed and ballast readjusted, if
necessary, to return empty weight CG to
within allowable limits.
16 DEC 2002
1
BHT-407-FMS-3
TC APPROVED
Section 2
NORMAL PROCEDURES
2
2-7. BEFORE TAKEOFF
1. P A R T S E P p u r g e s w it c h – A s
required.
2-7-A. BEFORE FLIGHT WHEN
OPERATING IN SNOW
CONDITIONS
2. Check engine air plenum chamber
through plexiglass windows on each
side of inlet cowling for snow, slush,
or ice, paying particular attention to
firewalls and rear face of particle
separator. Clean thoroughly before
each flight.
2-9. IN-FLIGHT OPERATIONS
1. T h o r o u g h l y c h e c k c a b i n r o o f ,
transmission cowling, deflector
baffles and engine air intake areas.
All areas checked shall be clean and
free of accumulated snow, slush, and
ice before each flight.
1. P A R T S E P p u r g e s w it c h – A s
required.
2-10. DESCENT AND LANDING
1. P A R T S E P p u r g e s w it c h – A s
required.
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3
No change from basic manual.
Section 4
PERFORMANCE
4
4-2. POWER ASSURANCE
CHECK
P e r f o r m a n c e i s r e d u c e d w i t h pa r t i c l e
separator installed. This reduction increases
2
16 DEC 2002
with use of particle separator purge and is
primarily result of bleed air being taken from
engine. A Power assurance check chart
(Figure 4-1) is provided to determine if engine
can produce installed power.
TC APPROVED
BHT-407-FMS-3
4
EXAMPLE: ENTER CHART AT OBSERVED TORQUE (70%)
PROCEED VERTICALLY DOWN TO PRESSURE ALTITUDE (6,000 FT.)
FOLLOW HORIZONTALLY TO THE RIGHT TO OBSERVED OAT (10 DEG.C)
DROP DOWN TO READ MAXIMUM ALLOWABLE MGT (681 DEG.C)
PARTICLE SEPARATOR KIT
MODEL 407 POWER ASSURANCE CHECK - ROLLS ROYCE 250-C47B ENGINE
HOVER OR LEVEL FLIGHT (85 TO 105 KIAS - NOT TO EXCEED VNE)
ENGINE TORQUE - PERCENT
35
40
45
50
55
60
65
70
75
80
85
90
95
100
175
200
225
250
275
300
325
350
375
400
425
450
475
500
PARTICLE SEPARATOR PURGE OFF
GENERATOR LOAD 35 AMPS OR LESS
POWER TURBINE - 100% RPM
HEATER / ECS OFF
ANTI-ICE OFF
138
128
118
108
98
88
78
68
58
48
525
550
575
600
625
650
675
700
725
750
775
MEASURED GAS TEMPERATURE - DEG.C
407FMS3 FIG 4-1 REV A.WMF
Figure 4-1. Power assurance check
16 DEC 2002
3
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
00
REFER TO
FIG. 4-5B
22
16
,0
26
MAX DEMONSTRATED HD
AREA B
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
0
,0
0
12
,0
00
0
20
T
AT
OA
00
°C
0
-1
10
AT
0
-2
10
UM
XI M
UM O
80
O
0
-3
MA
MAXIM
60
00
30
0
-4
MINIMUM OAT
24
17,000 FT HD
00
18
,0
20
,0
00
20,000 FT HD
-20
-2
00
0
A
SE
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
LE
VE
L
20
T
-F
P
H
-40
7
50 51.
00
40
00
40
REFER TO
FIG. 4-5
60
GROSS WEIGHT - LBS x 100
M407_FMS-3_FIG 4-2 (1_OF_4).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 1 of 4)
4
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
0
-3
12
00
,0
00
0
-2
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
AT
80
°C
10
0
-1
MINIMUM OAT
AT
0 0
UM O
MAXIM
O
5
-4
,0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-2_(2_OF_4).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 2 of 4)
16 DEC 2002
5
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
00
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
00
,0
10
00
°C
20
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
AT
80
-
10
MINIMUM OAT
AT
UM O
MAXIM
12
,0
0
0
14,000 FT HD
O
0 0
2 0 -1
30 0 -4
,0
00
REFER TO
FIG. 4-5A
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-2_(3_OF_4).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 3 of 4)
6
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
°C
0
,0
0
12
00
,0
10
00
-
80
AT
AT
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
5
MINIMUM OAT
O
0 0 0
0 -2 - 1
0 -3
-4
UM O
MAXIM
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-2_(4_OF_4).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 4 of 4)
16 DEC 2002
7
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
17,000 FT HD
16
14
00
,0
10
00
80
40
40
00
60
30
00
20
T
OA
AT
00
LE
V
EL
20
REFER TO
FIG. 4-5
SE
-2
00
0
A
T
-F
P
H
50
5 1 .7
-20
°C
10
M
MINIMUM OAT
-
U
XIM
UM O
-40
AT
MA
MAXIM
12
,0
00
14,000 FT HD
O
20 -10 0
30 0 -4
,0
00
REFER TO
FIG. 4-5A
AREA B
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-3_(1_OF_4).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 1 of 4)
8
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
°C
0
,0
0
12
00
,0
10
00
-
80
AT
AT
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
5
MINIMUM OAT
O
0 0
0 -2 -1 0
0 -3
-4
UM O
MAXIM
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-3_(2_OF_4).WMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 2 of 4)
16 DEC 2002
9
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
12
,0
00
0
0 -20 -1
0 -3
-4
00
80
AT
O
-
AT
0
UM O
10
,0
00
MAXIM
60
10
00
°C
MINIMUM OAT
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
EL
20
00
40
00
20
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-3_(3_OF_4).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 3 of 4)
10
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
00
,0
14
AT
00
,0
10
80
00
MINIMUM OAT
O
AT
0
-1
UM O
0
0 -2
-3
MAXIM
12
0
-4
,0
00
14,000 FT HD
EXTERNAL = 6000 LB
16
REFER TO
FIG. 4-5A
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
,0
18
,0
00
20
,0
00
20,000 FT HD
60
°C
00
-
0
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
EL
20
00
40
00
5
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-3_(4_OF_4).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 4 of 4)
16 DEC 2002
11
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-3__FIG_4-4_(1_OF_4).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 1 of 4)
12
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
18
20
22
24
00
REFER TO
FIG. 4-5A
26
AREA B
00
0
14,000 FT HD
°C
,0
00
12
00
0
-2
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
T
80
-
0
-3
0
AT
-4
10
O
5
M OA
MINIMUM OAT
0
U
MAXIM
,0
0
0
0
-1
14
,
16
,0
16
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FM S- 3__F IG_4-2_(2_OF_4).WMF
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 2 of 4)
16 DEC 2002
13
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-3__FIG_4-4_(3_OF_4).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 3 of 4)
14
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-3__FIG_4-4_(4_OF_4).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 4 of 4)
16 DEC 2002
15
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-3__FIG_4-5_(1_OF_4).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 1 of 4)
16
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
18
20
22
24
26
00
REFER TO
FIG. 4-5A
00
0
14,000 FT HD
-
00
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
5
°C
T
80
AT
O
0
M OA
MINIMUM OAT
10
,0
00
12
U
MAXIM
,0
0
0
0 10
0 -2 0 -3
-4
14
,
16
,0
16
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-3__FIG_4-5_(2_OF_4).WMF
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 2 of 4)
16 DEC 2002
17
BHT-407-FMS-3
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-3__FIG_4-5_(3_OF_4).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 3 of 4)
18
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-3__FIG_4-5_(4_OF_4).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 4 of 4)
16 DEC 2002
19
BHT-407-FMS-3
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
14,500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 70 FT/MIN FOR PURGE ON
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
-10 -20
6
0
OA
T°C
10
MA
16000
5
X
20
T
LIM
4
IT
PRESSURE ALTITUDE - FEET
OA
14000
12000
30
10000
3
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
18000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(1_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 1 of 10)
20
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
10,500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 70 FT/MIN FOR PURGE ON
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
1
2
3
5
X
MA
20000
4
IT
IM
TL
16000
PRESSURE ALTITUDE - FEET
OA
OA
18000
6
7
8
9
10
-2
0
T°C
-3
0
6
-1
0
0
5
10
14000
4
20
12000
10000
3
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
0
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(2_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 2 of 10)
16 DEC 2002
21
BHT-407-FMS-3
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
11,000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 60 FT/MIN FOR PURGE ON
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
-4
0
18000
-1
-2
0
0
OA
T°C
0
14000
10
MA
4
XO
12000
AT
20
LIM
3
IT
10000
30
8000
2
6000
40
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
-30
4000
1
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(3_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 3 of 10)
22
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
7500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 60 FT/MIN FOR PURGE ON
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
1
2
3
4
5
6
7
8
9
6
X
MA
-4
0
OA
18000
TL
PRESSURE ALTITUDE - FEET
-2
0
IT
IM
16000
10
OA
-3
0
5
-1
0
T°C
14000
0
4
12000
10
10000
3
20
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
0
20000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(4_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 4 of 10)
16 DEC 2002
23
BHT-407-FMS-3
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 170 FT/MIN ABOVE
8500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 55 FT/MIN FOR PURGE ON
Gross Weight 4000 lb (1814 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
18000
-2
0
-3
0
-1
14000
MA
OA
T-
0
4
10
IT
IM
TL
OA
10000
0
°C
X
12000
-4
0
3
20
8000
30
2
6000
4000
40
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
10
1
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(5_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 5 of 10)
24
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 170 FT/MIN ABOVE
4500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 55 FT/MIN FOR PURGE ON
Gross Weight 4000 lb (1814 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
M
AX
18000
O
M
LI
IT
PRESSURE ALTITUDE - FEET
-4
0
-3
0
OA
T°C
5
-2
0
14000
-1
0
4
0
12000
10
10000
3
20
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
AT
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(6_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 6 of 10)
16 DEC 2002
25
BHT-407-FMS-3
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 150 FT/MIN ABOVE
6000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 50 FT/MIN FOR PURGE ON
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
20000
6
7
8
9
10
6
-40
18000
-2 0
OA
T
12000
10000
-1
-°C
3
10
AT
XO
20
IT
L IM
8000
4
0
0
MA
PRESSURE ALTITUDE - FEET
-3
0
14000
2
6000
30
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(7_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 7 of 10)
26
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 150 FT/MIN ABOVE
1500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 50 FT/MIN FOR PURGE ON
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
16000
PRESSURE ALTITUDE - FEET
-2
0
14000
OA
T-
°C
5
-3
0
-1
12000
0
4
0
0
10000
10
8000
3
20
2
6000
PRESSURE ALTITUDE - 1000 METERS
-4
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(8_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 8 of 10)
16 DEC 2002
27
BHT-407-FMS-3
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
4500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
18000
-40
-20
14000
-1
12000
OA
10000
-3
0
4
0
0
T°C
3
MA
10
X
OA
8000
TL
20
2
IT
IM
6000
30
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
10
4000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(9_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 9 of 10)
28
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
14000
-4
-3
OA
T°C
12000
-1
0
10000
0
4
0
-2
0
3
0
10
8000
20
2
6000
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-6_(10_OF_10).WMF
Figure 4-6. Rate of climb takeoff power (sheet 10 of 10)
16 DEC 2002
29
BHT-407-FMS-3
TC APPROVED
407FMS-3-4-7-1.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 1 of 10)
30
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 225 FT/MIN ABOVE
5500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 70 FT/MIN FOR PURGE ON
Gross Weight 3000 lb (1361 kg)
RATE OF CLIMB - M/SEC
1
3
4
5
6
7
8
6
-3
AT
-2
I
LIM
0
0
5
T
16000
10
-4
0
XO
18000
-1
A
O
0
°C
T-
PRESSURE ALTITUDE - FEET
9
MA
20000
2
14000
0
12000
4
10
10000
3
20
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
0
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-7_(2_OF_10).WMF
Figure 4-7. Rate of climb maximum continuous power (sheet 2 of 10)
16 DEC 2002
31
BHT-407-FMS-3
TC APPROVED
407FMS-3-4-7-3.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 3 of 10)
32
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
RATE OF CLIMB
60 KIAS
HEATER ON
PARTICLE SEPARATOR
PURGE OFF
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 190 FT/MIN ABOVE
2500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 60 FT/MIN FOR PURGE ON
Gross Weight 3500 lb (1587 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
MA
XO
18000
AT
LI M
C
-1
0
12000
-2
0
4
0
10
10000
20
3
8000
2
6000
PRESSURE ALTITUDE - 1000 METERS
°
T-
14000
0
-3
0
A
O
PRESSURE ALTITUDE - FEET
5
-4
IT
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-3__FIG_4-7_(4_OF_10).WMF
Figure 4-7. Rate of climb maximum continuous power (sheet 4 of 10)
16 DEC 2002
33
BHT-407-FMS-3
TC APPROVED
407FMS-3-4-7-5.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 5 of 10)
34
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
407FMS-3-4-7-6.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 6 of 10)
16 DEC 2002
35
BHT-407-FMS-3
TC APPROVED
407FMS-3-4-7-7.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 7 of 10)
36
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
407FMS-3-4-7-8.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 8 of 10)
16 DEC 2002
37
BHT-407-FMS-3
TC APPROVED
407FMS-3-4-7-9.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 9 of 10)
38
16 DEC 2002
TC APPROVED
BHT-407-FMS-3
407FMS-3-4-7-10.TIF
Figure 4-7. Rate of climb maximum continuous power (sheet 10 of 10)
16 DEC 2002
39/40
MANUFACTURER’S DATA
BHT-407-FMS-3
Section 1
SYSTEMS DESCRIPTION
1
This kit is equipped with a PART SEP switch
located on overhead console. When switch is
OFF, engine bleed air is not used to purge
debris from particle separator. When switch is
ON, engine bleed is used to purge debris.
16 DEC 2002
41/42
BHT-407-FMS-4
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
SNOW DEFLECTOR
206-706-208
CERTIFIED
1 MARCH 1996
This supplement shall be attached to Model 407 Flight
Manual when Snow Deflector 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
COPYRIGHT NOTICE
2002
COPYRIGHT
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD. ALL RIGHTS
RESERVED
POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101
REISSUED — 16 DECEMBER 2002
BHT-407-FMS-4
NOTICE PAGE
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
16 DEC 2002
BHT-407-FMS-4
LOG OF REVISIONS
Original ....................... 0....................01 MAR 96
Revision ...................... 1....................27 MAY 97
Revision ...................... 2....................04 SEP 98
Revision ......................3 ................... 17 APR 00
Reissue........................0 ................... 16 DEC 02
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
FLIGHT MANUAL
Title............................................................ 0
NP .............................................................. 0
A — B ........................................................ 0
C/D .............................................................0
i/ii ...............................................................0
1 — 76 ........................................................0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
16 DEC 2002
A
BHT-407-FMS-4
TC APPROVED
LOG OF TC APPROVED REVISIONS
Original ........................0 ...................01 MAR 96
Revision.......................1 ...................27 MAY 97
Revision.......................2 ...................04 SEP 98
B
16 DEC 2002
Revision ...................... 3 ................... 17 APR 00
Reissue ....................... 0.................... 16 DEC 02
BHT-407-FMS-4
LOG OF FAA APPROVED REVISIONS
Original ....................... 0................... 01 MAR 96
Revision ...................... 1................... 27 MAY 97
Revision ...................... 2................... 18 MAY 99
Revision ......................3 ................... 05 MAY 00
Reissue........................0 .................... 24 SEP 03
16 DEC 2002
C/D
BHT-407-FMS-4
GENERAL INFORMATION
Snow deflector kit (206-706-208) consists of
two deflectors that mount on either side of
transmission fairing, just forward of engine
air inlets.
16 DEC 2002
i/ii
TC APPROVED
BHT-407-FMS-4
Section 1
LIMITATIONS
1
1-3. TYPES OF OPERATION
Snow deflector kit shall be installed for
operation in falling or blowing snow. They
may be installed with basic inlet screen or
particle separator kit (206-706-212).
1-6. WEIGHT AND CENTER OF
GRAVITY
Actual weight changes shall be determined
after kit is installed and ballast readjusted, if
necessary, to return empty weight CG to within
allowable limits. Refer to Center of gravity vs
weight empty chart in BHT-407-MM-1.
1-11. AMBIENT
TEMPERATURES
Snow deflectors shall be removed for
operations above 30°C (86°F).
1-22. SNOW OPERATION
For operation in falling or blowing snow, the
following limits apply:
Hover flight in falling and/or blowing
snow is limited to 15 minute duration
after which helicopter shall be landed
and checked for snow and/or ice
accumulation.
Flight operations are prohibited when
visibility in falling or blowing snow is
less than one-half (1/2) statute mile.
16 DEC 2002
1
BHT-407-FMS-4
TC APPROVED
Section 2
NORMAL PROCEDURES
2
2-3. PREFLIGHT CHECK
2. Particle separator kit (if installed),
check engine air plenum chamber
through plexiglass windows on each
side of inlet cowling for snow, slush,
or ice, paying particular attention to
firewalls and rear face of particle
separator. Clean thoroughly before
each flight.
2-3-A. EXTERIOR CHECK
2-3-A-1.
OPERATION IN FALLING OR
BLOWING SNOW
1. Thoroughly check cabin roof,
transmission fairing, deflector
baffles, and engine air intake areas.
All areas checked shall be clean and
free of accumulated snow, slush,
and ice before each flight.
2-3-A-2.
AFTER EXITING HELICOPTER
WARNING
FAILURE TO INSTALL ENGINE
IN T A K E C O V E R S C O U L D A L L O W
FALLING/BLOWING SNOW TO ENTER
THE PARTICLE SEPARATOR PLENUM
(IF INSTALLED).
NOTE
Due to reduced performance at higher
temperatures, it is recommended that
snow deflectors be removed above
20°C (68°F).
Install protective covers (engine intake,
exhaust, and pitot tube).
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3
No change from basic manual.
2
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
Section 4
PERFORMANCE
4
4-1. INTRODUCTION
4-5. HOVER CEILING
Refer to appropriate performance charts in
accordance with optional equipment installed.
4-5-A. HOVER CEILING INGROUND-EFFECT
NOTE
Due to reduced performance at higher
temperatures, it is recommended that
snow deflectors be removed above
20°C (68°F).
4-2. POWER ASSURANCE
CHECK
Power assurance check chart Figure 4-1.
Performance is reduced with snow baffles
installed. Power assurance check chart
(Figure 4-1) is provided to determine if engine
c a n p r o d u c e i n s ta l l e d p o w e r. P o w e r
assurance check shall be conducted in level
flight only. This chart is valid for both basic
inlet with snow deflectors installed and snow
deflectors with particle separator installed.
PARTICLE SEP PRG switch (if installed) shall
be OFF when performing a power assurance
check.
Hover ceiling IGE (takeoff power) charts are
presented in Figure 4-2, and Hover ceiling IGE
(maximum continuous power) charts are
presented in Figure 4-3.
4-5-B. HOVER CEILING OUT-OFGROUND-EFFECT
Hover ceiling OGE (takeoff power) charts are
presented in Figure 4-4, and Hover ceiling
OGE (maximum continuous power) charts are
presented in Figure 4-5.
4-7. CLIMB AND DESCENT
4-7-A. RATE OF CLIMB
Rate of climb (takeoff power) charts are
presented in Figure 4-6, Rate of climb
(maximum continuous power) charts are
presented in Figure 4-7, Rate of climb (takeoff
p o w e r ) ( pa r t i c l e s e pa r a t o r ) c h a r ts a r e
presented in Figure 4-8, and Rate of climb
(maximum continuous power) (particle
separator) charts are presented in Figure 4-9.
16 DEC 2002
3
BHT-407-FMS-4
TC APPROVED
EXAMPLE: ENTER CHART AT OBSERVED TORQUE (70%)
PROCEED VERTICALLY DOWN TO PRESSURE ALTITUDE (6,000 FT.)
FOLLOW HORIZONTALLY TO THE RIGHT TO OBSERVED OAT (10 DEG.C)
DROP DOWN TO READ MAXIMUM ALLOWABLE MGT (721 DEG.C)
SNOW DEFLECTOR KIT
MODEL 407 POWER ASSURANCE CHECK - ROLLS ROYCE 250-C47B ENGINE
PARTICLE SEPARATOR PURGE OFF
GENERATOR LOAD 35 AMPS OR LESS
POWER TURBINE - 100% RPM
HEATER / ECS OFF
ANTI-ICE OFF
LEVEL FLIGHT (85 TO 105 KIAS - NOT TO EXCEED VNE)
ENGINE TORQUE - PERCENT
35
40
45
50
55
60
65
70
75
80
85
90
95
100
200
225
250
275
300
325
350
375
400
425
450
475
500
128
118
108
98
88
78
68
58
48
175
525
550
575
600
625
650
675
700
725
750
775
MEASURED GAS TEMPERATURE - DEG.C
407FMS4 FIG 4-1.WMF
Figure 4-1. Power assurance check chart
4
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
00
16
,
00
0
18
,0
20
,0
00
14
0
-4
AT
REFER TO
FIG. 4-5A
AREA B
MO
17,000 FT HD
MAX DEMONSTRATED HD
U
XIM
REFER TO
FIG. 4-5B
MA
,0
00
20,000 FT HD
12
00
,0
0
-2
30
0
-1
00
°C
0
10
-3
20
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
AT
80
AT
UM O
MAXIM
MINIMUM OAT
O
10
,0
0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(1_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 1 of 8)
16 DEC 2002
5
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
00
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
0
-1
,0
00
REFER TO
FIG. 4-5A
12
0
00
,0
-2
0
00
°C
10
-
5
0
-4
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
AT
80
AT
-3
UM O
MAXIM
MINIMUM OAT
O
0
,0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(2_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 2 of 8)
6
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
-4
,0
00
REFER TO
FIG. 4-5A
0
12
,0
0
00
,0
10
00
°C
80
-
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
20
00
AT
10
MINIMUM OAT
0
UM O
MAXIM
AT
O
0 -1 0
30 -2
0 -
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(3_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 3 of 8)
16 DEC 2002
7
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
00
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
°C
00
,0
10
00
80
-
MINIMUM OAT
AT
O
AT
0
0 20 -1 0
-3 -
UM O
MAXIM
12
0
-4
,0
0
0
14,000 FT HD
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
5
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(4_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 4 of 8)
8
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
,0
00
,0
00
14
0
-4
00
,0
0
-2
30
00
°C
0
-3
10
-
20
-1
AT
80
AT
O
10
12
,0
0
0
0
14,000 FT HD
UM O
MAXIM
MINIMUM OAT
26
AREA B
T
REFER TO
FIG. 4-5A
24
MAX DEMONSTRATED HD
OA
16
,
00
0
17,000 FT HD
22
UM
18
,0
20
20
X IM
REFER TO
FIG. 4-5B
18
MA
00
20,000 FT HD
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
0
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(5_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 5 of 8)
16 DEC 2002
9
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
00
,0
10
00
80
0
°C
-2
10
0 -4
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
0
-3
MINIMUM OAT
AT
O
5
AT
0
UM O
MAXIM
12
,0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(6_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 6 of 8)
10
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
12
,0
0
0
0
0 -3
-4
°C
00
,0
10
00
-
80
AT
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
20
00
AT
10
MINIMUM OAT
O
0 0 0
-2 - 1
UM O
MAXIM
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(7_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 7 of 8)
16 DEC 2002
11
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
12
00
,0
10
00
°C
80
AT
AT
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
5
00
0
MINIMUM OAT
O
UM O
MAXIM
0 0
0 -2 -1
0 -3
-4
,0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-2_(8_OF_8).WMF
Figure 4-2. Hover ceiling IGE – takeoff power (sheet 8 of 8)
12
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
17,000 FT HD
16
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
00
,0
12
00
,0
00
°C
80
-
-20
LE
V
REFER TO
FIG. 4-5
SE
-2
00
0
A
T
-F
P
H
-40
EL
20
00
40
00
60
30
00
20
AT
10
UM O
10
AT
MAXIM
MINIMUM OAT
AREA B
O
0 0 0
0 -2 -1
0 -3
-4
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(1_OF_8).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 1 of 8)
16 DEC 2002
13
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
°C
00
,0
10
00
-
80
AT
MINIMUM OAT
O
AT
0 2 0 10 0
-3 - -
UM O
MAXIM
12
0
-4
,0
0
0
14,000 FT HD
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
-2
00
SE
P
H
-F
A
T
LE
VE
L
20
00
40
00
60
00
5
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(2_OF_8).WMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 2 of 8)
14
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
14
,0
00
REFER TO
FIG. 4-5A
12
,0
00
14,000 FT HD
-
00
,0
10
00
80
AT
AT
O
UM O
0 30 2 0 0
-4 - - -1
MAXIM
60
00
0
°C
MINIMUM OAT
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
00
10
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
EL
20
00
40
20
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(3_OF_8).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 3 of 8)
16 DEC 2002
15
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
14
,0
00
REFER TO
FIG. 4-5A
12
,0
00
14,000 FT HD
00
80
AT
-
60
O
00
AT
0 0 0 0
-4 -3 -2 -1
UM O
10
,0
00
MAXIM
MINIMUM OAT
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
00
°C
LE
V
EL
20
SE
-2
00
0
A
T
-F
P
H
-20
5
00
40
0
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(4_OF_8).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 4 of 8)
16
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
17,000 FT HD
16
14
,0
00
REFER TO
FIG. 4-5A
14,000 FT HD
00
,0
12
°C
00
-
10
60
20
00
AT
80
AT
UM O
10
,0
00
MAXIM
MINIMUM OAT
AREA B
O
0 0 0
0 -2 -1
0 -3
-4
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
EL
20
00
40
00
30
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(5_OF_8).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 5 of 8)
16 DEC 2002
17
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
00
MAX DEMONSTRATED HD
AREA B
17,000 FT HD
16
,
00
0
18
,0
20
,0
00
20,000 FT HD
14
,0
00
REFER TO
FIG. 4-5A
°C
12
00
,0
10
00
80
-
AT
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
5
00
0
MINIMUM OAT
AT
O
UM O
MAXIM
0 0
0 -2 -1
0 -3
-4
,0
0
0
14,000 FT HD
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(6_OF_8).WMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 6 of 8)
18
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
MAX DEMONSTRATED HD
00
,0
14
12
,0
00
14,000 FT HD
AT
00
,0
00
80
AT
O
0 0 20 0
- 4 -3 - - 1
UM O
10
EXTERNAL = 6000 LB
16
REFER TO
FIG. 4-5A
MAXIM
MINIMUM OAT
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
,0
18
,0
00
20
,0
00
20,000 FT HD
00
-
60
°C
0
LE
V
EL
20
SE
-2
00
0
A
T
-F
P
H
-20
20
00
40
00
10
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(7_OF_8).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 7 of 8)
16 DEC 2002
19
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
16
18
20
22
24
26
REFER TO
FIG. 4-5B
16
14
,0
00
REFER TO
FIG. 4-5A
12
,0
00
14,000 FT HD
00
,0
10
AT
80
00
UM O
0 0 0
-4 -3 -2
MAXIM
-
60
00
0
-1
AT
O
MINIMUM OAT
AREA B
MAX GW
INTERNAL = 5000 LB
00
17,000 FT HD
EXTERNAL = 6000 LB
MAX DEMONSTRATED HD
,0
18
,0
00
20
,0
00
20,000 FT HD
°C
LE
V
SE
-2
00
0
A
T
-F
P
H
-20
EL
20
00
5
40
00
0
-40
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-3_(8_OF_8).EMF
Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 8 of 8)
20
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
00
REFER TO
FIG. 4-5A
22
24
26
AREA B
-
0
,0
0
,0
00
12
00
0
00
SE
-2
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
T
80
1
0 -
°C
-2
10
AT
10
20
30
M OA
MINIMUM OAT
O
0
-3
U
MAXIM
0
-4
0
14
,
00
0
20
0
AT
14,000 FT HD
18
XO
MA
16
,0
16
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-4_(1_OF_8).WMF
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 1 of 8)
16 DEC 2002
21
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
18
20
22
24
00
REFER TO
FIG. 4-5A
26
AREA B
00
0
14,000 FT HD
°C
-4
5
0
00
,0
00
-
10
AT
0
0
-2
-20
0
-2
00
SE
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
VE
LE
A
T
-F
P
H
-40
MAX GW
INTERNAL = 5000 LB
REFER TO
FIG. 4-5
L
20
00
40
00
60
00
T
80
O
0
-3
12
0
M OA
MINIMUM OAT
-1
U
MAXIM
,0
0
0
14
,
16
,0
16
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-4_(2_OF_8).WMF
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 2 of 8)
22
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-4_(3_OF_8).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 3 of 8)
16 DEC 2002
23
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-4_(4_OF_8).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 4 of 8)
24
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
00
22
24
26
AREA B
10
,0
00
12
T
80
00
MINIMUM OAT
M OA
30
°C
20
A T 30 O
-
1 0 20
U
MAXIM
,0
0
0
0
-4
14
,
0
AT
14,000 FT HD
00
0
20
XO
REFER TO
FIG. 4-5A
18
MA
16
,0
16
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
60
00
0
-1
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-4_(5_OF_8).WMF
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 5 of 8)
16 DEC 2002
25
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-4_(SHEET_6).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 6 of 8)
26
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-4_(SHEET_7).EPS
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 7 of 8)
16 DEC 2002
27
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
18
20
22
24
26
00
REFER TO
FIG. 4-5A
00
0
14,000 FT HD
,0
0
10
,0
00
12
00
80
°C
00
AT
60
O
T
0 0
-2 -1 0
M OA
MINIMUM OAT
0
-3
U
MAXIM
0
-4
0
14
,
16
,0
16
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
5
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-4_(8_OF_8).WMF
Figure 4-4. Hover ceiling OGE – takeoff power (sheet 8 of 8)
28
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-5_(SHEET_1).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 1 of 8)
16 DEC 2002
29
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
17,000 FT HD
18
20
22
24
26
00
REFER TO
FIG. 4-5A
00
0
14,000 FT HD
,0
00
12
10
0
-
00
60
°C
00
AT
80
O
T
0 10 0
-2 -
M OA
MINIMUM OAT
-3
U
MAXIM
,0
0
0
-4
0
14
,
16
,0
16
-40
-20
0
20
OAT - °C
40
60 32
36
40
EXTERNAL = 6000 LB
MAX GW
INTERNAL = 5000 LB
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
40
00
5
REFER TO
FIG. 4-5
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-5_(2_OF_8).WMF
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 2 of 8)
30
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-5_(SHEET_3).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 3 of 8)
16 DEC 2002
31
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW DEFLECTOR
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-5_(SHEET_4).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 4 of 8)
32
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-5_(SHEET_5).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 5 of 8)
16 DEC 2002
33
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-5_(SHEET_6).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 6 of 8)
34
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
REFER TO
FIG. 4-5A
REFER TO FIG. 4-5
M407_FMS-4__FIG_4-5_(SHEET_7).EPS
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 7 of 8)
16 DEC 2002
35
BHT-407-FMS-4
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW DEFLECTOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART
IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)
GROSS WEIGHT - KG x 100
20
00
REFER TO
FIG. 4-5A
00
0
14,000 FT HD
0
,0
0
0
-4
10
00
T
0
-2
M OA
80
26
0
,0
00
12
24
-3
U
MAXIM
0
AT
60
0
00
-1
O
MINIMUM OAT
22
EXTERNAL = 6000 LB
18
14
,
16
,0
16
MAX GW
INTERNAL = 5000 LB
17,000 FT HD
°C
40
00
-
-40
-20
0
00
-2
H
P
SE
-F
A
T
LE
VE
L
20
00
5
REFER TO
FIG. 4-5
0
20
OAT - °C
40
60 32
36
40
44
48
52
56
60
GROSS WEIGHT - LBS x 100
M407_FMS-4__FIG_4-5_(8_OF_8).WMF
Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 8 of 8)
36
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-6-1
Figure 4-6. Rate of climb (takeoff power) (sheet 1 of 10)
16 DEC 2002
37
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-6-2
Figure 4-6. Rate of climb (takeoff power) (sheet 2 of 10)
38
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-6-3
Figure 4-6. Rate of climb (takeoff power) (sheet 3 of 10)
16 DEC 2002
39
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-6-4
Figure 4-6. Rate of climb (takeoff power) (sheet 4 of 10)
40
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-6-5
Figure 4-6. Rate of climb (takeoff power) (sheet 5 of 10)
16 DEC 2002
41
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-6-6
Figure 4-6. Rate of climb (takeoff power) (sheet 6 of 10)
42
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-6-7
Figure 4-6. Rate of climb (takeoff power) (sheet 7 of 10)
16 DEC 2002
43
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-6-8
Figure 4-6. Rate of climb (takeoff power) (sheet 8 of 10)
44
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
SNOW DEFLECTOR
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
2500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
18000
-40
14000
-2
M
AX
LI O
M AT
IT
12000
-1
OA
T°C
0
-3
0
4
0
0
10000
10
8000
20
3
2
6000
30
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
10
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-4__FIG_4-6_(9_OF_10).WMF
Figure 4-6. Rate of climb (takeoff power) (sheet 9 of 10)
16 DEC 2002
45
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-6-10
Figure 4-6. Rate of climb (takeoff power) (sheet 10 of 10)
46
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-7-1
Figure 4-7. Rate of climb (maximum continuous power) (sheet 1 of 10)
16 DEC 2002
47
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-7-2
Figure 4-7. Rate of climb (maximum continuous power) (sheet 2 of 10)
48
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-7-3
Figure 4-7. Rate of climb (maximum continuous power) (sheet 3 of 10)
16 DEC 2002
49
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-7-4
Figure 4-7. Rate of climb (maximum continuous power) (sheet 4 of 10)
50
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-7-5
Figure 4-7. Rate of climb (maximum continuous power) (sheet 5 of 10)
16 DEC 2002
51
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-7-6
Figure 4-7. Rate of climb (maximum continuous power) (sheet 6 of 10)
52
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
RATE OF CLIMB
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
SNOW DEFLECTOR
REDUCE RATE OF CLIMB 150 FT/MIN ABOVE
1500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 4500 lb (2041 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
-4
14000
12000
10000
-2
AT
X O IT
M A L IM
O
AT
-1
-°C
0
0
4
0
0
10
3
20
8000
30
2
6000
-4 0
4000
1
-20
2000
0
PRESSURE ALTITUDE - FEET
-3
5
0
PRESSURE ALTITUDE - 1000 METERS
16000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-4__FIG_4-7_(7_OF_10).WMF
Figure 4-7. Rate of climb (maximum continuous power) (sheet 7 of 10)
16 DEC 2002
53
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-7-8
Figure 4-7. Rate of climb (maximum continuous power) (sheet 8 of 10)
54
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-7-9
Figure 4-7. Rate of climb (maximum continuous power) (sheet 9 of 10)
16 DEC 2002
55
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-7-10
Figure 4-7. Rate of climb (maximum continuous power) (sheet 10 of 10)
56
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-8-1
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 1 of 10)
16 DEC 2002
57
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-8-2
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 2 of 10)
58
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-8-3
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 3 of 10)
16 DEC 2002
59
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-8-4
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 4 of 10)
60
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-8-5
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 5 of 10)
16 DEC 2002
61
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-8-6
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 6 of 10)
62
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-8-7
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 7 of 10)
16 DEC 2002
63
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-8-8
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 8 of 10)
64
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR - PURGE OFF
SNOW DEFLECTOR
REDUCE RATE OF CLIMB 135 FT/MIN ABOVE
2000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5000 lb (2268 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
18000
-4 0
14000
-2
M
AX
12000 LI O
M AT
IT
OA
-1
T-
°C
0
4
0
-3
0
0
10
10000
3
20
8000
30
2
6000
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
10
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-4__FIG_4-8_(9_OF_10).WMF
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 9 of 10)
16 DEC 2002
65
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-8-10
Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 10 of 10)
66
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-9-1
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 1 of 10)
16 DEC 2002
67
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-9-2
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 2 of 10)
68
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-9-3
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 3 of 10)
16 DEC 2002
69
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-9-4
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 4 of 10)
70
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-9-5
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 5 of 10)
16 DEC 2002
71
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-9-6
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 6 of 10)
72
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-9-7
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 7 of 10)
16 DEC 2002
73
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-9-8
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 8 of 10)
74
16 DEC 2002
TC APPROVED
BHT-407-FMS-4
407FMS-4-4-9-9
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 9 of 10)
16 DEC 2002
75
BHT-407-FMS-4
TC APPROVED
407FMS-4-4-9-10
Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 10 of 10)
76
16 DEC 2002
BHT-407-FMS-5
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
CARGO HOOK
206-706-341
AND
407-704-023
CERTIFIED
14 FEBRUARY 1996
This supplement shall be attached to the BHT-407-FM-1
when the Cargo Hook kit has been installed.
Information contained herein supplements 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
COPYRIGHT
2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
REISSUE — 15 NOVEMBER 2007
REVISION 1 — 25 MARCH 2008
BHT-407-FMS-5
NOTICE PAGE
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 or maintenance 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———15 NOV 2007
BHT-407-FMS-5
LOG OF REVISIONS
Original ..................... 0...................... 14 FEB 96
Revision .................... 1...................... 04 SEP 98
Reissue......................0 ..................... 15 NOV 07
Revision.....................1 .....................25 MAR 08
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
FLIGHT MANUAL
Title.............................................................1
NP ...............................................................0
A/B..............................................................1
C/D..............................................................1
E/F ..............................................................1
i/ii ................................................................1
1 ..................................................................0
2 ..................................................................1
3 – 10 ..........................................................0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
25 MAR 2008—Rev. 1———A/B
BHT-407-FMS-5
LOG OF TC APPROVED REVISIONS
Original ..................... 0...................... 14 FEB 96
Revision .................... 1...................... 04 SEP 98
APPROVED
Reissue...................... 0 ..................... 15 NOV 07
Revision .................... 1 ..................... 25 MAR 08
DATE
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
25 MAR 2008—Rev. 1———C/D
BHT-407-FMS-5
LOG OF FAA APPROVED REVISIONS
Original ..................... 0...................... 14 FEB 96
Revision .................... 1......................18 MAY 99
Reissue...................... 0 ..................... 21 DEC 07
Revision .................... 1 ..................... 05 MAY 08
25 MAR 2008—Rev. 1———E/F
BHT-407-FMS-5
GENERAL INFORMATION
Installation of Cargo Hook kit (206-706-341) or Cargo Hook Retrofit kit (407-704-023) with an
approved hook assembly adds capability of transporting external cargo. Kit contains electrical
and manual releases, both operated from the pilot seat. Cargo hook is located at FS 121.0 (3073
mm).
Approved assemblies may have a spring-loaded keeper (P/N 17149-9 or 528-010-ALL) or a
keeperless hook (P/N 528-023-ALL) design.
Cargo hook kit will permit operator to use helicopter for transportation of external cargo.
25 MAR 2008—Rev. 1———i/ii
TC APPROVED
BHT-407-FMS-5
Section 1
LIMITATIONS
1-3.
TYPES OF OPERATION
Operation of helicopter with no load on
external cargo suspension hook is authorized
under standard airworthiness certificate
without removing unit from helicopter.
With a load attached to suspension assembly,
operations shall be conducted in accordance
with appropriate operating rules for external
loads.
1-6.
WEIGHT AND CENTER OF
GRAVITY
Actual weight change shall de determined
after cargo hook is installed and ballast
readjusted, if necessary, to return empty
weight CG to within allowable limits. Refer to
Center of Gravity vs Weight Empty Chart in
the BHT-407-MM-1.
CAUTION
LOADS THAT RESULT IN GROSS
WEIGHTS ABOVE 5000 POUNDS
(2268 KG) SHALL BE CARRIED ON
CARGO HOOK AND SHALL BE
JETTISONABLE.
Maximum gross weight of helicopter and
external load operations is 6000 pounds
(2724 kg).
1
Maximum cargo hook load is 2650 pounds
(1202 kg).
Refer to the BHT-407-FM-1 for Gross Weight
Center of Gravity Limits charts for external
cargo operations.
1-7.
AIRSPEED
VNE with external cargo load is 100 KIAS.
CAUTION
AIRSPEED WITH EXTERNAL CARGO
IS LIMITED BY CONTROLLABILITY.
CAUTION SHOULD BE EXERCISED
WHEN CARRYING EXTERNAL
CARGO,
AS
HANDLING
CHARACTERISTICS MAY BE
AFFECTED BY SIZE, WEIGHT, AND
SHAPE OF CARGO LOAD.
Light weight, high drag loads require a swivel
connector between cargo hook and sling to
prevent unstable oscillations in flight above
20 KIAS.
1-20. INSTRUMENT MARKINGS
AND PLACARDS
Refer to Figure 1-3 for decals.
15 NOV 2007———1
BHT-407-FMS-5
TC APPROVED
Kit 206-706-341-109
Location: On cargo hook roller beam.
Kits 206-706-341-141 and 407-704-023-103
Location: On cargo hook beam.
407-FMS-5_0001
Figure 1-3. Placards and Decals
2———Rev. 1—25 MAR 2008
TC APPROVED
BHT-407-FMS-5
Section 2
NORMAL PROCEDURES
2-2.
2
FLIGHT PLANNING
2-2-A.
RELEASE IF LOAD RING IS TOO
LARGE.
GROUND CREW INSTRUCTIONS
Instruct ground crew member to discharge
helicopter static electricity before attaching
cargo by touching airframe with ground wire,
or, if metal sling is used, hookup ring can be
struck against cargo hook. If contact has been
lost after initial grounding, helicopter shall be
electrically regrounded and, if possible,
c o n ta c t m a i n t a i n e d u n t i l h o o k u p i s
completed.
1. Cargo hook — Condition and security.
Instruct ground personnel to check
primary load ring and secondary load
ring for condition and proper size
(Table 2-1). Check for correct rigging
(Figure 2-1 and Figure 2-2).
WARNING
USE OF INAPPROPRIATELY SIZED
LOAD RINGS MAY RESULT IN LOAD
HANG-UP WHEN LOAD RING IS TOO
SMALL OR INADVERTENT LOAD
CAUTION
THE EXAMPLES SHOWN IN
Figure 2-1 THROUGH Figure 2-3 ARE
NOT INTENDED TO REPRESENT
ALL POSSIBILITIES. IT IS THE
RESPONSIBILITY
OF
THE
OPERATOR TO MAKE SURE THE
HOOK WILL FUNCTION PROPERLY
WITH THE RIGGING. SOME
COMBINATIONS OF SMALL
PRIMARY RINGS AND LARGE
SECONDARY RINGS COULD CAUSE
FOULING DURING RELEASE.
2. Check that only one primary ring is
captured in the load beam and only
one secondary ring with correct
cross-section dimension is captured
in the primary ring. Additional rings,
slings, or shackles shall be attached
to the secondary load ring (Figure 2-1
through Figure 2-3).
Table 2-1: Cargo Hook Ring Sizes
Primary Ring
Inside Diameter
Primary Ring
Cross-section
Secondary Ring
Maximum Cross-section
CARGO HOOK P/N 17149-6
1.50 to 1.68 inches
(38.10 to 42.67 mm)
0.75 inch
(19.05 mm)
0.438 inch
(11.12 mm)
CARGO HOOK P/N 528-010-ALL
1.50 to 1.87 inches
(38.10 to 47.50 mm)
0.875 inch
(22.22 mm)
0.625 inch
(15.88 mm)
CARGO HOOK P/N 528-023-ALL (KEEPERLESS HOOK)
Refer to Figure 2-4.
15 NOV 2007———3
BHT-407-FMS-5
2-3.
PREFLIGHT CHECK
2-3-B.
TC APPROVED
2-8.
1. Hover helicopter at sufficient height to
allow ground crew member to
discharge static electricity and attach
cargo sling to cargo hook.
EXTERIOR CHECK
Cargo suspension assembly — Condition and
security.
Cargo sling — Condition, proper length.
2-4.
2-4-A.
NOTE
For keeperless hook, prior to
external load attachment, ensure
hook is in OPEN position by
m o m e n ta r ily d ep r e ss i ng c y cl ic
CARGO RELEASE switch. Ground
personnel must ensure that load
beam is latched closed after load is
applied.
INTERIOR AND PRESTART
CHECK
INTERIOR CHECK
(SPRING-LOADED KEEPER)
1. CARGO HOOK circuit breaker — In.
2. Cyclic CARGO RELEASE switch —
Press and release; pull down on cargo
hook; hook should open. Release
cargo hook; hook should close and
lock.
2. Ascend vertically, directly over load,
then slowly lift load from surface.
3. Pedals — Check
directional control.
3. EMERG CARGO RELEASE PULL
handle — Pull and hold; pull down on
cargo hook; hook should open. Push
handle in; hook should close and lock.
2-4-B.
INTERIOR CHECK (KEEPERLESS
HOOK)
1. Cargo hook circuit breaker — In.
2. Cyclic CARGO RELEASE switch —
Press and release; hook should open.
Push hook up manually and verify that
it latches closed.
3. EMERG CARGO RELEASE PULL
handle — Pull and hold; pull down on
cargo hook; hook should open. Push
handle in; push hook up manually and
verify that it latches closed.
2-7.
BEFORE TAKEOFF
CARGO HOOK circuit breaker — In.
EMERG CARGO RELEASE PULL handle — In.
4———15 NOV 2007
TAKEOFF
for
adequate
4. Hover power — Check TORQUE
required to hover with external load.
5. Take off into wind, if possible,
allowing
adequate
sling
load
clearance over obstacles.
2-9.
IN-FLIGHT OPERATIONS
NOTE
Control movements should be made
smoothly and kept to a minimum to
prevent oscillation of sling load.
EMERG CARGO RELEASE PULL
handle will function regardless of
CARGO RELEASE switch position.
1. AIRSPEED — Within limits for
adequate controllability of helicopter
load combination.
2. Flight path — As planned to avoid
flight with external load over any
person, vehicle, or structure.
TC APPROVED
2-10. DESCENT AND LANDING
1. Flight path and approach angle — As
required for wind direction and
obstacle clearance.
BHT-407-FMS-5
2. Execute approach to a hover with load
clear of surface. When stabilized at a
hover, descend slowly until load
contacts surface. Maintain tension on
sling.
3. Cyclic CARGO RELEASE switch —
Press to release sling from hook.
15 NOV 2007———5
BHT-407-FMS-5
TC APPROVED
Figure 2-1. External Load Rigging (Cargo Hook P/N 17149-6)
6———15 NOV 2007
TC APPROVED
BHT-407-FMS-5
CORRECT RIGGING
1 MAXIMUM CROSS SECTION
OF PRIMARY RING
PRIMARY 1
RING I.D.
1
MAXIMUM CROSS SECTION
OF SECONDARY RING
LOAD
INCORRECT RIGGING
INCORRECT RIGGING
MULTIPLE RINGS
ON LOAD BEAM
MULTIPLE RINGS
ON LOAD BEAM
NOTE
1
Refer to Table 2-1.
407_FMS_5_0002
Figure 2-2. External Load Rigging (Cargo Hook P/N 528-010-ALL)
15 NOV 2007———7
BHT-407-FMS-5
TC APPROVED
PRIMARY RING 1
SECONDARY RING
LOAD
NOTE
1
Refer to Figure 2-4.
407_FMS_5_0004
Figure 2-3. External Load Rigging (Cargo Hook P/N 528-023-ALL)
8———15 NOV 2007
TC APPROVED
BHT-407-FMS-5
1.97 IN. (50 mm)
MIN.
1.02 IN. (26 mm)
MAX.
2.16 IN. (55 mm)
MIN.
PRIMARY RING
SECONDARY RING
407_FMS_5_0003
Figure 2-4. Cargo Hook P/N 528-023-ALL Ring Sizes
15 NOV 2007———9
BHT-407-FMS-5
TC APPROVED
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3-11. CARGO FAILS TO RELEASE
ELECTRICALLY
3
REGARDLESS OF CARGO RELEASE
SWITCH POSITION.
In the event the cargo hook will not release
the sling when the cyclic CARGO RELEASE
switch is pressed, proceed as follows:
WARNING
EMERG CARGO RELEASE PULL
HANDLE
WILL
FUNCTION
1. Maintain tension on sling.
2. Pull EMERG CARGO RELEASE PULL
handle to release load.
Section 4
PERFORMANCE
4-5.
4
HOVER CEILING
Refer to the BHT-407-FM-1 for out of ground
effect hover performance.
There is no change from the BHT-407-FM-1
performance with no load attached to the
cargo hook.
Performance may be affected by size and
shape of external load.
Section 5
WEIGHT AND BALANCE
5-2.
EMPTY WEIGHT CENTER OF
GRAVITY
Load on hook is at FS 121.0 (3073 mm).
10———15 NOV 2007
5
B HT-407-FMS-6
407
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
AUXILIARY FUEL KIT
407-706-011
CERTIFiED
20 MARCH 1996
This supplement shall be attached to Model 407
Flight Manual when AUXILIARY FUEL KIT 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
BELL HELICOPTER INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVION OF TEXTRON CANADA LTD.
Bell Helicopterk1*1i.1Tl
A Subsidiary 01 Textron inc.
POST OFFICE BOX 482 • FORT WORTH, TEXAS 18101
20 MARCH 1 996
BHT-407-FMS-6
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-407-FMS-6
LOG OF REVISIONS
Original
.0 . 20 MAR 96
LOG OF PAGES
REViSION
NO.
PAGE
FLIGHT MANUAL
Title—NP
A—B
PAGE
C/D
i/il
0
0
1—6
REVISION
NO.
0
0
0
NOTE
Revised text is indicated by a black vertical line, insert latest revision pages; dispose of
superseded pages.
A
BHT-407-FMS-6
DOT APPROVED
LOG OF APPROVED REViSIONS
Original
.0 . 20 MAR 96
APPROVED:
-CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
B
DATE: Z0
p1&& 6
BHT-407-FMS-6
LOG OF FAA APPROVED REVISIONS
Original
.0 . 20 MAR 96
C/D
DOT APPROVED
BHT-407-FMS-6
GENERAL INFORMATION
Auxiliary fuel kit (407-706-011) consists of fuel tank, tubing, electrical wiring, micro-switch
and hardware for installation. Removable 19 U.S. gallons (71.9 liters) fuel tank is mounted
in baggage compartment to aft bulkhead. Fuel transfers between main aft fuel cell and
auxiliary fuel cell for filling and emptying by gravity.
i/u
BHT-407-FMS-6
DOT APPROVED
Section 1
LIMITA TIONS
1-6. WEIGHT AND CENTER OF
GRAVITY
if necessary, to return empty weight CG to
within allowable limits. Refer to Center of
gravity vs weight empty chart in BHT-407MM-i.
Actual weight changes shall be determined
after kit is installed and ballast readjusted,
L Section 2
NORMAL PROCEDURES
2-2. FLIGHT PLANNING
2-3. PREFLIGHT CHECK
With auxiliary fuel tank installed, full fuel
Baggage compartment — Check auxiliary
fuel tank security and condition.
indication is approximately 1005 lbs. (146.9
U.S. gallons/556 liters) of Jet A.
DOT APPROVED
BHT-407-FMS-6
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
No change from basic manual.
Section 4
PERFORMA NCE
No change from basic manual.
2
BHT-407-FMS-6
Section 5
WEIGHT AND BALANCE
5-7. FUEL LOADING
Longitudinal center of gravity of fuel shifts
as it is consumed (Figure
5-1). Extreme
Figure 5-1
effects of fuel consumption on helicopter
center of gravity for standard fuel system
2. Critical fuel for computing most aft
useful load is 146.9 U.S. gallons (556
liters).
figure 5-2
Fuel loading tables (figure
5-2) list usable
fuel quantities, weight, and moments in
both U.S. and metric units.
are as follows:
1. Critical fuel for computing most
forward useful load is 74.5 U.S.
gallons (282.0 liters).
3
BHT-407-FMS-6
160———--————--———————--——--———-——
150 — — — — — — — — — — — — — — — — — — — — — — — —
:
11111
11111 11111;
F
120
/- -——- -s———
110
U)
0
-———-—.-
-
( 90
Cu
—
-———-
-———-
80
70
U60
50
40
30
20
10
0
114
116
118
120
122
124
126
128
130
132
134
136
136
Fuselage Station-Inches
600
550
500
450
400
U) 350
a)
300
V
- —--------z——------— ——-—
—
—
V -—--—
—————— ——
————— ——
a,
CL. 250
200
150
--—
--—-
—.---_
100
50
0
2900
3000
3100
3200
3300
-
3400
Fuselage Station-Millimeters
Figure 5-1. Fuel Center of Gravity - Auxiliary Fuel.
4
3500
407-FS-5-t
BHT-407-FMS-6
AUXILIARY FUEL LOADING (U.S.)
QUANTITY WE1GHT
FUEL LOADING TABLE (U.S.)
LONGITUDINAL
JP-5, JP-8
C.G.
MOMENT
QUANTITY
WEIGHT
(U.S. GAL)
(IN)
JP-4
(LBS)
(IN-LBS)
(U.S. GAL)
4345
5
10
8,775
15
13,250
5
32.5
133.7
10
15
20
65.0
97.5
130.0
135.0
135.9
136.4
25
162.5
136.7
22,214
28.4
184.6
137.0
25,290
30
195.0
35
40
227.5
260.0
134.3
127.8
122.9
26,189
29,075
31,954
45
292.5
119.1
34,837
50
325.0
328.9
357.5
390.0
422.5
455.0
116.0
115.7
37,70
484.3
487.5
520.0
552.5
585.0
617.5
650.0
682.5
116.1
116.3
118.0
119.6
121.0
122.3
123.4
124.5
38,054
41,506
45,318
49,095
52,826
56,227
56,696
61,360
66,079
70,785
75,520
80,210
84,971
* 50.6
55
60
65
70
074.5
75
80
85
90
95
116.1
116.2
116.2
116.1
17,732
(LBS)
34.0
68.0
102.0
136.0
170.0
LONGITUDINAL
C.G.
MOMENT
408.0
442.0
476.0
506.6
510.0
544.0
578.0
612.0
646.0
680.0
714.0
748.0
782.0
116.2
116.2
123.4
124.5
125.5
126.5
(IN-LBS)
4546
9,180
13,862
18,550
23,239
26,455
27,397
30,416
33,429
36,445
39,440
39,812
43,421
47,410
51,360
55,264
58,816
59,313
64,192
69,129
74,052
79,006
83,912
88,893
93,874
98,923
(IN)
193.1
133.7
135.0
135.9
136.4
136.7
137.0
204.0
238.0
272.0
134.3
127.8
122.9
45
50
306.0
119.1
340.0
* 50.6
344.1
374.0
116.0
115.7
20
25
28.4
30
35
40
55
60
65
70
074.5
75
80
85
90
95
116.1
116.1
116.1
116.3
118.0
119.6
121.0
122.3
715.0
125.5
89,733
115
747.5
126.5
94,559
100
105
110
115
120
125
780.0
127.5
99,450
120
816.0
127.5
104,040
812.5
128.5
130
135
845.0
877.5
129.4
130.2
125
130
850.0
884.0
140
145
910.0
942.5
954.9
131.0
131.7
132.0
104,406
109,343
114,251
119,210
124,127
126,047
135
140
145
918.0
952.0
986.0
998.9
128.5
129.4
130.2
131.0
109,225
114,390
119,524
124,712
129,856
131,855
100
105
110
A 146.9
A 146.9
131.7
132.0
* MOST FORWARD FUEL C.G.
D CRITICAL FUEL FOR MOST FORWARD C.G. CONDITION
A FULL FUEL— CRITICAL FUEL FOR MOST AFT C.G. CONDITION
Figure 5-2. Auxiliary Fuel Loading - Sheet 1 of 2
5
BHT-407-FMS-6
AUXILIARY FUEL LOADING (METRIC)
FUEL LOADING TABLE (METRIC)
LONGITUDINAL
JP-5, JP-8
C.G.
MOMENT
QUANTITY
WEIGHT
JP-4
QUANTITY WEIGHT
(LITERS)
15
30
45
60
(mm)
(kg)
(kg-mm/100)
397
(LITERS)
15
799
30
24.4
1204
1613
45
36.7
3439
1262
60
75
90
48.9
3455
3465
3472
3478
3479
3352
3228
3129
3049
2982
1689
2117
2545
2977
3048
3278
23.4
35.0
46.7
3389
3415
3439
3455
75
58.4
3465
2024
90
105
107.5
120
135
150
165
180
70.1
3472
2434
81.8
3478
3479
3352
3228
3129
3049
2982
2938
2940
2949
2845
2912
3134
3393
3655
3918
4181
4383
*191.6
195
210
225
240
255
270
D282.0
285
300
315
330
345
360
375
390
405
420
435
450
465
480
495
510
525
540
555
A556.1
11.7
83.7
93.5
105.1
116.8
128.5
140.2
149.2
151.9
163.6
175.2
186.9
198.6
210.3
219.7
222.0
233.7
245.3
257.0
268.7
280.4
292.1
303.8
315.4
327.1
338.8
350.5
362.2
373.9
385.5
397.2
408.9
420.6
432.3
433.1
2951
2953
2950
2948
2949
2956
2991
3024
3054
3082
3107
3130
3152
3172
3192
3213
3234
3253
3272
3290
3306
3322
3337
3351
3352
LONGITUDINAL
C.G.
MOMENT
(mm)
(kg-mm/100)
3389
413
3415
833
4466
4825
5170
5519
5859
6200
6479
6562
6990
7418
7849
8281
8712
9143
9576
10004
10441
10886
11335
11782
12234
12683
13131
13584
14035
14486
14518
(kg)
12.2
61.1
73.3
85.6
87.6
97.8
105
107.5
120
135
180
110.0
122.2
134.4
146.7
* 191.6
156.1
195
210
225
240
255
270
158.9
150
165
UIJ 282.0
285
300
315
330
345
360
375
390
405
420
435
450
465
480
495
510
525
540
555
A556.1
171.1
183.3
195.6
207.8
220.0
229.8
232.2
244.5
256.7
2951
2953
2950
2948
2949
2956
2991
3024
268.9
3054
281.1
293.3
305.6
3082
3107
317.8
330.0
342.2
354.5
366.7
378.9
391.1
403.3
415.6
427.8
440.0
452.2
453.1
* MOST FORWARD FUEL C.G.
U CRITICAL FUEL FOR MOST FORWARD C.G. CONDITION
A FULL FUEL— CRITICAL FUEL FOR MOST AFT C.G. CONDITION
Figure 5-2. Auxiliary Fuel Loading - Sheet 2 of 2
6
2938
2940
2949
3130
3152
3172
3192
3213
3234
3253
3272
3290
3306
3322
3337
3351
3352
3551
3824
4098
4375
4586
4672
5046
5409
5776
6130
6486
6777
6864
7313
7763
8212
8664
9113
9565
10017
10468
10923
11390
11859
12326
12797
13269
13740
14212
14683
15153
15188
BHT-407-FMS-7
32If 4107
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
LITTER(S)
407-706-631
OR
407-799-100
AND
407-799-001
CERTIFIED
14 FEBRUARY 1996
This supplement shall be attached to Model 407 Flight
Manual when LITTER(S) 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
19
COPYRIGHT
BELL® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell Helicopter TEXTRON
A SubsidIary of Textron Inc.
POST OFFICE BOX 412 • FORT WORTH, TEXAS 71101
25 SEPTEMBER 1997
REVISION 1 —16 SEPTEMBER 1999
BHT-407-FMS-7
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 761 01-0482
NP
BHT-407-FMS-7
LOG OF REVISIONS
Original
Revision
Reissue
Reissue
.0.14 Feb 96
1
25 Apr 96
0
0
09 MAY 97
25 SEP 97
RevisIon
.1
.16 SEP 99
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
CID
FLIGHT MANUAL
i/il
Title
1
NP
0
A—B
1
A
1I2.
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
Rev.1
A
BHT-407-FMS-7
LOG OF DOT APPROVED REVISIONS
Original
Revision
Reissue
.0. 14 Feb 96
1
0
25 Apr 96
09 MAY 97
0
Reissue
Revision
1
DATE
CHIEF, FLIGHT TEST
FOR DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
B
Rev.1
25 SEP 97
16 SEP 99
BHT-407-FMS-7
LOG OF FAA APPROVED REVISIONS
Original
Revision
Reissue
0
14 Feb 96
1
25 Apr 96
09 MAY 97
0
Reissue
Revision
0
25 SEP 97
24 SEP 99
1
Rev.1
C/D
BHT-407-FMS-7
GENERAL INFORMATION
Litter kit (407-706-631) provides helicopter with capability to carry one patient on litter
with room and access for medical attendants. The principal configuration consists of two
parts, basic provisions kit and litter assembly kit. Basic provisions contain structural
brackets and all necessary hardware for supporting a litter. Litter assembly contains a
folding aluminum litter with patient restraints. In addition, an optional injured skier
provisions kit is available. In this configuration, horizontal support bar located behind copilot seat is moveable and may be secured with quick release pins in either normal or an
upper location. This feature provides an additional 6 inches (15.24 cm) of clearance above
patient when support bar is installed in upper position. Basic litter provisions with litter
assembly adds 27 pounds (12.3 kilograms) to empty weight of helicopter. Injured skier
provisions kit adds an additional 1.5 pounds (0.7 kilograms) to empty weight.
Customized litter kit (407-799-100) is same as basic litter kit installation with injured skier
provisions, except that support bar is bolted in place which may be desired for a
permanent EMS configured helicopter.
Dual Litter Kit (407-799-001) provides helicopter with capability to carry two litter patients.
This kit contains structural supports and all necessary hardware to install a second litter
above standard litter. Basic litter kit provisions (407-706-631) must be installed in
conjunction with this kit. If helicopter is equipped with injured skier provisions kit, dual
litter kit allows upper litter patient to be placed in elevated foot position.
i/u
BHT-407-FMS-7
DOT APPROVED
Section 1
LIMITA TIONS
1-5. CONFIGURATION
Copilot cyclic and collective controls shall
be removed and stowed
installed.
when litter is
Patient(s) shall be restrained by litter
straps.
STRUCTURAL SUPPORT MUST BE
INSTALLED IN THE UPPER POSITION OR
LOWER POSITION FOR FLIGHT.
Location: On copilot seat back support
assembly and on forward side of vertical
tunnel.
This placard applicable only with Auxiliary
Litter Kit with Injured Skier Provisions
installed.
1-6. WEIGHT AND CENTER OF
GRAVITY
Actual weight change shall be determined
after kit is installed and ballast readjusted,
if necessary, to return empty weight CG to
within allowable limits. Refer to Center of
gravity vs weight empty chart in BHT-407MM-i.
CO-PILOT SEAT SHALL
NOT BE OCCUPIED UNLESS
PROTECTIVE COVERS ARE
INSTALLED ON UPPER
LITTER SUPPORT BRACKETS
TYPICAL
Location: On forward side of interior trim
panel centered on door post between
upper and lower litter support brackets.
1-20. INSTRUMENT
MARKINGS AND PLACARDS
This placard applicable with basic Litter
I Kit installed.
1/2
BHT-407-FMS-1 7
13211407
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
CARGO TIE-DOWN
PROVISIONS KIT
407-705-201
CERTIFIED
1 APRIL 1996
This supplement shall be attached to Model 407
Flight Manual when CARGO TlEDOWN
PROVISIONS KIT 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
1996
COPYRIGHT
BELL HELICOPTER INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell Helicopterki *41L1.ii
A Subsidiary of Textron Inc.
OFFICE BOX 482 • FORT WORTH, TEXAS 76101
1 APRIL
BHT-407-FMS-1 7
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-407-FMS-17
LOG OF REVISIONS
Original
.0 . 1 APR 96
LOG OF PAGES
REVISION
NO.
PAGE
C/D
FLIGHT MANUAL
Title—NP
A—B
PAGE
i/u
0
0
1/2
REVISION
NO.
0
0
0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
A
DOT APPROVED
BHT-407-FMS-17
LOG OF APPROVED REVISIONS
Original
.0
.1 APR 96
DATE:
AP,9Y,
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
B
/ "'pp
BUT-407-FMS-17
LOG OF FAA APPROVED REVISIONS
Original
0
1 APR 96
C/D
BHT-407-FMS-17
GENERAL INFORMATION
Cargo tie-down provisions kit (407-705-201) provides forward bulkhead tie-down
provisions using four (4) shackle/eyebolt assemblies and floor mounted provisions using
four (4) anchor plates. These provisions allow cargo to be secured with a tie-down
assembly.
i/il
BHT-407-FMS-17
DOT APPROVED
Section I
LIMITA TIONS
1-20. INSTRUMENT
MARKINGS AND PLACARDS
NO CARGO ABOVE THIS LINE
W.L. 55
CAUTION
WHEN AFT FACING SEAT AREA IS
USED FOR CARGO:
-DO NOT REMOVE SEAT CUSHIONS.
-MAXIMUM ALLOWABLE CARGO
WEIGHT - 100 LBS.
-CARGO MUST BE SECURED TO PREVENT
IN FLIGHT MOVEMENT.
-CARGO WEIGHT TO BE UNIFORMLY
DISTRIBUTED.
Location: Each side of center post.
Location: Each side of center post.
1/2
BHT-407-FMS-20
'!407
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
KLN 89B GPS NAViGATOR
407-705-001
CERTIFIED
14 FEBRUARY 1996
This supplement shall be attached to Model 407 Flight
Manual when KLN 89B GPS NAVIGATOR 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
1996
COPYRIGHT
BELL HELICOPTER INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell HelicopterkE *iit•1
A Subsidiary of Textron Inc.
POST OFFICE BOX 482 • FORT WORTH, TEXAS 16101
14 FEBRUARY 1996
REVISION 1 —26 NOVEMBER 1996
BHT-407-FMS-20
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-407-FMS-20
LOG OF REVISIONS
Original
Revision
.0
1
.14 FEB 96
26 NOV 96
LOG OF PAGES
REVISION
NO.
PAGE
FLIGHT MANUAL
Title
NP
1
0
REVISION
PAGE
NO.
A—B
1
CID
1
i/u
1
1—2
1
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
Rev.1
A
BHT-407-FMS-20
DOT APPROVED
LOG OF APPROVED REVISIONS
Original ........... 0
Revision
1
14 FEB 96
26 NOV 96
APPROVED:
DATE:
',z
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION BRANCH
DEPARTMENT OF TRANSPORT
B
Rev.1
BHT-407-FMS-20
LOG OF FAA APPROVED REVISIONS
Original
Revision
.0
1
.14 FEB 96
26 NOV 96
Rev.1
C/D
BHT-407-FMS-20
GENERAL INFORMATiON
The KLN 89B GPS Navigator is a
navigator's aid for use in ICAO defined
unit. If GPS is coupled to the KCS 55A
gyrocompass with Kl 525A HSI Kit 407-705-
worldwide geographic regions as defined
in the King KLN Pilots Guide.
002 the system will additionally include a
The system consists of a combined GPS
receiver and navigational computer, an
Visual navigation data, when selected, is
presented on the pilot HSI in the form of Li
R steering, bearing-to-waypoint and TO!
antenna, and associated wiring. Visual
Navigation data is presented on the GPS
NAV/GPS Switch/Annunicator.
FROM indications.
Rev. 1
i/u
I
BHT-407-FMS-20
DOT APPROVED
Section 1
LIM1TA TIONS
1-1. INTRODUCTION
2.
A KLN 89B Pilots Guide (King p/n 00608786-0000, Operational Revision Status
01) shall be accessible by the flight crew
at all times during flight.
It is the responsibility of the pilot
to verify that any navigation data
used is correct.
1-20. INSTRUMELNI
MARKINGS AND PLACARDS
The GPS navigator shall be operated in
accordance with the manufactures
instruction with the following exceptions:
f GPS LIMITED TO VFR USE ONLY
1. There is no air data or fuel
management data available in
this installation.
Lsection 2
NORMAL PROCEDURES
2-3. PREFLIGHT CHECK
2-7. BEFORE TAKEOFF
2-3-A. CABIN TOP
2-7-A. GPS
GPS antenna — Condition and security.
GPS unit — Turn on, verify operational
revision status on initial page is identical
2-4. INTERIOR AND
PRESTART CHECK
2-4-A. PRESTART CHECK
GPS and CAUTION LIGHTS circuit
breakers — In.
to that of available KLN 89B Pilot's Guide.
Pilots HSI course pointer (if installed) —
Align to desired course shown on GPS
display.
NAVIGPS switch-annunciator (if installed)
— Press, verify GPS segment illuminated
and NAV segment extinguished.
GPU unit — Verify off.
I
Rev.1
1
BHT-407-FMS-20
DOT APPROVED
NOTE
For additional normal procedures,
Pilot HSI deviation bar (if installed)
except air data and fuel
displayed.
89B Pilot's Guide.
Verify centered and TO indication
management data, refer to KLN
Section 3
EMERGENCWMALFUNCTION PROCEDURES
31. INTRODUCTION
NOTE
If GPS navigation system
becomes inoperative, continue
basic VFR navigation procedures.
2
Rev.1
BHT-407-FMS-21
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
FIRE DETECTION SYSTEM
407-799-004
OR
407-706-015
OR
407-706-025
CERTIFIED
2 MAY 1996
This supplement shall be attached to Model 407 Flight
Manual when FIRE DETECTION SYSTEM has been installed.
Information contained herein supplements 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
COPYRIGHT
2004
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
REISSUE — 07 JULY 2004
BHT-407-FMS-21
NOTICE PAGE
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 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———07 JUL 2004
BHT-407-FMS-21
LOG OF REVISIONS
Original ..................... 0......................02 MAY 96
Reissue ..................... 0...................... 08 SEP 98
Reissue......................0 ...................... 07 JUL 04
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
FLIGHT MANUAL
Title.............................................................0
NP ...............................................................0
A/B..............................................................0
C/D..............................................................0
E/F ..............................................................0
i/ii ................................................................0
1 — 2 ..........................................................0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
07 JUL 2004———A/B
BHT-407-FMS-21
LOG OF TC APPROVED REVISIONS
Original ..................... 0......................02 MAY 96
Reissue ..................... 0...................... 08 SEP 98
APPROVED
Reissue...................... 0 ...................... 07 JUL 04
DATE
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
07 JUL 2004———C/D
BHT-407-FMS-21
LOG OF FAA APPROVED REVISIONS
Original ..................... 0......................02 MAY 96
Reissue ..................... 0...................... 08 SEP 98
Reissue...................... 0 ...................... 02 SEP 04
07 JUL 2004———E/F
TC APPROVED
BHT-407-FMS-21
GENERAL INFORMATION
Bell Fire Detection System (407-799-004, 407-706-015 or 407-706-025) will illuminate ENGINE FIRE
warning light on instrument panel if an excessive temperature or fire develops in engine
compartment.
07 JUL 2004———i/ii
TC APPROVED
BHT-407-FMS-21
Section 1
LIMITATIONS
1-6. WEIGHT AND CENTER OF
GRAVITY
1
readjusted, if necessary, to return empty
weight CG to within allowable limits.
Actual weight change shall be determined
after system is installed and ballast
Section 2
NORMAL PROCEDURES
2-4. INTERIOR
CHECK
AND
PRESTART
2
FIRE DET TEST switch — Press, ENGINE FIRE
light illuminates, release, ENGINE FIRE light
extinguishes.
07 JUL 2004———1
BHT-407-FMS-21
TC APPROVED
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3
Table 3-1.
PANEL WORDING
FAULT CONDITION
CORRECTIVE ACTION
ENGINE FIRE
Excessive temperature
condition in engine
compartment
Immediately enter autorotation.
Throttle — Close.
FUEL VALVE switch — OFF.
If time permits, FUEL BOOST/XFR
circuit breaker switches — OFF
Execute a normal autorotation and
landing
BATT switch — OFF.
NOTE
Do not restart engine until cause of fire has been determined and corrected.
2———07 JUL 2004
B HT-407-FMS-22
'!407
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
AUXILIARY VERTICAL FIN
STROBE LIGHTS
407-899-023
CERTIFIED
10 MAY 1996
This supplement shall be attached to Model 407
Flight Manual when AUXILIARY VERTICAL FIN
STROBE LIGHTS 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.
________________
COPYRIGHT NOTICE
1996
COPYRIGHT
BELL HELICOPTER INC
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell Helicopterki 1 t.lfl
A Subsidiary of Textron Inc.
POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101
1 MA
BHT-407-FMS-22
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-407-FMS-22
LOG OF REVISIONS
Original
.0 . 10 May 96
LOG OF PAGES
PAGE
Title—NP
A—B
CID
REVISION
NO.
0
0
0
PAGE
i/u
1—2
REVISION
NO.
0
0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
A
DOT APPROVED
BHT-407-FMS-22
LOG OF APPROVED REVISIONS
Original ........... 0 . 10 May 96
APPROVED:
DATE:
'7a'y
o CHIEF, FLIGHT TEST
te4FoR
DIRECTOR — AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
B
6
BHT-407-FMS-22
LOG OF FAA APPROVED REVISIONS
Original ........... 0 . 10 May 96
do
BHT-407-FMS-22
GENERAL iNFORMATION
Auxiliary vertical fin strobe lights installation (407-899-023) consist of power supply unit
and two strobe lights installed on left and right auxiliary vertical fins.
i/u
DOT APPROVED
BHT-407-FMS-22
LII Section 1
LIMITA TIONS
1-5. CONFIGURATION
clouds or other weather
phenomena.
1-5-A. OPTIONAL EQUIPMENT
1-20. INSTRUMENT
MARKiNGS AND PLACARDS
Auxiliary vertical fin strobe lights are not
NIGHT OPERATION OF AUXILIARY VERTICAL
FIN STROBE LIGHTS IS PROHIBITED
approved for night operations.
NOTE
(Located on inst. panel - typical)
High intensity strobe lights
should not be used inflight when
there is an adverse reflection from
Lsection 2
NORMAL PROCEDURES
2-1. INTRODUCTION
FIN LT on/off CCT BKR/switch
located on overhead console.
NOTE
Both auxiliary vertical fin strobe
lights are controlled by AUX VERT
1
BHT-407-FMS-22
DOT APPROVED
Lsection 3
EMERGENCWMA LFUNCTION PROCEDURES
3-7. ELECTRICAL SYSTEM
NOTE
For emergency or malfunction
conditions, auxiliary vertical fin
2
strobe lights may be disabled by
selecting OFF at AUX VERT FIN
LT CCT BKR/switch. If auxiliary
vertical fin strobe lights become
inoperative, continue basic flight
procedures.
B HT-407- FM 5-23
'!407
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
RYAN TRAFFIC COLLISION
AVOIDANCE DEVICE
407-899-022
CERTIFIED
15 MAY1996
This supplement shall be attached to Model 407 Flight
Manual when RYAN TRAFFIC COLLISION AVOiDANCE
DEVICE ATS9000 has been installed in accordance with
407-899-022.
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
1996
COPYRIGHT
BELL HELICOPTER INC
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
1
I
Bell Helicopter ki* I
•1
A Subsidiary of Textron Inc.
POST OFFICE BOX 482 - FORT WORTH, TEXAS 78101
BHT-407-FMS-23
NOTICE PAGE
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. 0. Box 482
Fort Worth, Texas 76101-0482
NP
BHT-407-FMS-23
LOG OF REViSIONS
Original ........... 0 . 15 May 96
LOG OF PAGES
REVISION
PAGE
NO.
FLIGHT MANUAL
Title—NP
A—B
CID
i/u
1—2
REVISION
PAGE
NO.
0
0
314
5/6
0
0
0
0
0
MANUFACTURER'SDATA
7/8
0.
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
A
DOT APPROVED
BHT-407-FMS-23
LOG OF APPROVED REViSIONS
Original
.0 . 15 May 96
APPROVED:
CHIEF, FLIGHT TEST
FOR
DIRECTOR AIRWORTHINESS BRANCH
DEPARTMENT OF TRANSPORT
B
DATE:
BHT-407-FMS-23
LOG OF FAA APPROVED REVISIONS
Original
.0 . 15 May 96
C/D
BHT-407-FMS-23
GENERAL INFORMATiON
RYAN ATS9000 Traffic and Collision Avoidance Device (TCAD) (407-899-022) consists of
display unit, processor unit, transponder coupler, dual antenna module, two antennas,
wiring and hardware necessary for installation. A digital display is mounted in instrument
panel and contains all controls required to operate TCAD. Processor unit, transponder
coupler and dual antenna module are at various locations throughout helicopter,
depending on configuration. Antenna locations are on cabin top and underside of
helicopter.
RYAN TCAD is an on-board air traffic display used to identify potential collision threats.
TCAD computes relative altitude and distance of threats using transponder replies from
nearby Mode C equipped aircraft. Aircraft with non-Mode C transponders can provide
distance information. TCAD will not detect aircraft without operating transponders. Within
certain limits system creates a shield of airspace around helicopter, whereby detected
traffic cannot penetrate without generating an alert. Shield size is selectable for various
phases of flight and is adjustable by pilot.
Display is a bright, alphanumeric character. Distance is displayed in nautical miles (NM)
and relative altitude is displayed in 100 foot increments. TCAD is capable of displaying
multiple threats.
For familiarization of all ATS9000 TCAD features and operation, refer to Pilots Handbook,
P/N 32-21 02 Revision 1 or later.
i/il
DOT APPROVED
BHT-407-FMS-23
Section 1
LIMITA TIONS
1-6. WEIGHT AND CENTER OF
GRAVITY
if necessary, to return empty weight CG to
within allowable limits. Refer to Center of
gravity vs weight empty chart in BHT-407MM-i.
Actual weight change shall be determined
after kit is installed and ballast readjusted,
L Section 2
NORMAL PROCEDURES
2-4. INTERIOR AND
PRESTART CHECK
2-11. ENGINE SHUTDOWN
MIJTE/PWR button — Pull (off)
MUTE/PWR button — Push (on).
1
BHT-407-FMS-23
DOT APPROVED
Section 3
EMERGENC V/MALFUNCTION PROCEDURES
1. MUTE/PWR button — Pull (off)
3-7 ELECTRICAL SYSTEM
and continue flight.
3-7-A. TCAD MALFUNCTION
TCAD malfunction is annunciated by
words Signal Fail, SgnlFail, Link Failure or
Interface Fail displayed on TCAD display.
Table 3-1.
PANEL
WORDING
FAULT CONDITION
TRAFFIC
Proximate traffic detected.
Locate intruder aircraft using see
and avoid concept.
(advisory)
NOTE
TCAD is advisory only. Operations
shall be conducted in accordance
with operational regulations in
effect at helicopter location.
2
CORRECTIVE ACTION
DOT APPROVED
BHT-407-FMS-23
Section 4
PERFORMANCE
No change from basic manual.
3/4
BHT-407-FMS-23
Lsection 5
WEIGHT AND BALANCE
No change from basic manual.
5/6
MANUFACTURER'S DATA
BHT-407-FMS-23
L Section 1
SYSTEMS DESCRIPTION
: :.,.,
I TRAFFICj
tt:)
Figure 1-1. Caution and waring panel
7/8
BHT-407-FMS-25
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
QUIET CRUISE MODE
407-706-016
CERTIFIED
8 MAY 1998
This supplement shall be attached to the BHT-407-FM-1
when the Quiet Cruise Mode kit is installed.
Information contained herein supplements 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
COPYRIGHT
2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
REISSUE — 17 DECEMBER 2002
REVISION 1 — 30 JULY 2008
BHT-407-FMS-25
NOTICE PAGE
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 or maintenance 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. 1—30 JUL 2008
BHT-407-FMS-25
LOG OF REVISIONS
Original ..................... 0......................08 MAY 98
Reissue ..................... 0......................18 MAY 99
Reissue......................0 ..................... 17 DEC 02
Revision.....................1 ...................... 30 JUL 08
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
MANUFACTURER’S DATA
FLIGHT MANUAL
Title.............................................................1
NP ...............................................................1
A/B..............................................................1
C/D..............................................................1
E/F ..............................................................1
i/ii ................................................................1
1 – 7............................................................0
8 ..................................................................1
9 – 12 ..........................................................0
13/14 ...........................................................0
15/16 ...........................................................0
REVISION
NO.
17 – 18........................................................ 0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
30 JUL 2008—Rev. 1———A/B
BHT-407-FMS-25
LOG OF TC APPROVED REVISIONS
Original ..................... 0......................08 MAY 98
Reissue ..................... 0......................18 MAY 99
APPROVED
Reissue...................... 0 ..................... 17 DEC 02
Revision .................... 1 ...................... 30 JUL 08
DATE
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
30 JUL 2008—Rev. 1———C/D
BHT-407-FMS-25
LOG OF FAA APPROVED REVISIONS
Original ..................... 0....... Not FAA Approved
Reissue ..................... 0......................27 MAY 99
Reissue...................... 0 ...................... 24 SEP 03
Revision .................... 1 ..................... 18 AUG 08
30 JUL 2008—Rev. 1———E/F
BHT-407-FMS-25
GENERAL INFORMATION
Installation of Quiet Cruise Mode kit (407-706-016) permits flight operations at 92% NR when
above 50 KIAS and 200 feet AGL. Flyover noise level is reduced by 3.8 dBA SEL when in Quiet
Cruise Mode. Kit consists of electrical selector switch on collective, annunciator on instrument
panel, and additional markings on dual tachometer.
30 JUL 2008—Rev. 1———i/ii
TC APPROVED
BHT-407-FMS-25
Section 1
LIMITATIONS
1
1-3.
TYPES OF OPERATION
1-6-B.
Quiet cruise mode is approved for VFR
operations only.
1-5.
1-5-A.
CONFIGURATION
REQUIRED EQUIPMENT
FADEC system software 5.201 or higher is
required for Quiet Cruise Mode operations.
1-5-B.
OPTIONAL EQUIPMENT
Helicopters S/N 53000 – 53074 shall be in
compliance with Technical Bulletin 407-96-2
(Increase in VNE).
1-5-D.
CARGO HOOK
Cargo hook operations while in Quiet Cruise
Mode is not approved.
1-6.
WEIGHT AND CENTER
OF GRAVITY
Actual weight change shall be determined
after kit is installed and ballast readjusted, if
necessary, to return empty CG to within
allowable limits. Refer to Center of gravity vs
weight empty chart in BHT-407-MM-2.
1-6-A.
WEIGHT
Maximum GW for Quiet Cruise Mode
operation is 5000 pounds (2268 kilograms).
CENTER OF GRAVITY —
QUIET CRUISE MODE
OPERATION
For longitudinal CG limits refer to Gross
weight longitudinal center of gravity limits
chart (Figure 1-1).
For lateral CG limits refer to Gross weight
lateral center of gravity limits chart (Figure 12).
1-7.
1-7-A.
AIRSPEED
QUIET CRUISE MODE
NOTE
Refer to Section 4, HEIGHT – VELOCITY
ENVELOPE.
Minimum airspeed is 50 KIAS.
VNE is 100 KIAS.
1-8.
ALTITUDE
NOTE
Refer to Section 4, HEIGHT – VELOCITY
ENVELOPE.
Minimum altitude is approximately 200 feet
AGL.
Maximum altitude is 6,000 feet HD.
17 DEC 2002
1
BHT-407-FMS-25
TC APPROVED
1-13. POWER PLANT
1-15. ROTOR
1-13-B. POWER TURBINE RPM
(NP)
1-15-A. ROTOR RPM – POWER ON
1-13-B-1.
QUIET CRUISE MODE
Minimum
91.5%
Continuous operation
91.5 to 92.5%
Maximum continuous
92.5%
1-13-D. ENGINE TORQUE
Engine torque is restricted to maximum
continuous power (93.5%) while in Quiet
Cruise Mode.
2
17 DEC 2002
1-15-A-1.
QUIET CRUISE MODE
Continuous operation
91.5 to 92.5%
Maximum continuous
92.5%
1-20. INSTRUMENT
MARKINGS AND
PLACARDS
Refer to Figure 1-3 for Placards and decals.
Refer to Figure 1-4 for Instrument markings.
TC APPROVED
BHT-407-FMS-25
LONGITUDINAL C.G.
5400
5200
FS 127.0
5000
4800
4600
GROSS WEIGHT - POUNDS
4400
QUIET CRUISE MODE AFT LIMIT
4200
4000
3800
3600
3400
3200
3000
2800
2600
2400
118
119
120
121
122
123
124
125
126
127
128
129
130
FUSELAGE STATION - INCHES
M407_FMS-25__FIG_1-1_(1_OF_2).WMF
Figure 1-1. Gross weight longitudinal center of gravity limits (Sheet 1 of 2)
17 DEC 2002
3
BHT-407-FMS-25
TC APPROVED
LONGITUDINAL C.G.
2400
2300
2268
2200
2100
GROSS WEIGHT - KILOGRAMS
2000
QUIET CRUISE MODE AFT LIMIT
1900
1800
1700
1600
1500
1400
1300
1200
1100
3000
3226
3025
3050
3075
3100
3125
3150
3175
3200
3225
3250
3275
3300
FUSELAGE STATION - MILLIMETERS
M407_FMS-25__FIG_1-1_(2_OF_2).WMF
Figure 1-1. Gross weight longitudinal center of gravity limits (Sheet 2 of 2)
4
17 DEC 2002
TC APPROVED
BHT-407-FMS-25
LATERAL C.G.
5400
5200
-0.7
5000
0.9
4800
4600
GROSS WEIGHT - POUNDS
4400
4200
4000
3800
3600
-1.1
1.3
3400
QUIET CRUISE MODE LIMITS
3200
3000
2800
2600
2400
-3
-2
-1
0
1
2
3
BUTTOCK LINE - INCHES
M407_FMS-25__FIG_1-2_(1_OF_2).WMF
Figure 1-2.
Gross weight lateral center of gravity limits (Sheet 1 of 2)
17 DEC 2002
5
BHT-407-FMS-25
TC APPROVED
LATERAL C.G.
2400
2300
-18
23
2268
2200
2100
GROSS WEIGHT - KILOGRAMS
2000
1900
1800
1700
-28
1600
33
1500
QUIET CRUISE MODE LIMITS
1400
1300
1200
1100
-60
-50
-40
-30
-20
-10
0
10
20
30
40
50
60
BUTTOCK LINE - MILLIMETERS
M407_FMS-25__FIG_1-2_(2_OF_2).WMF
Figure 1-2.
6
17 DEC 2002
Gross weight lateral center of gravity limits (Sheet 2 of 2)
TC APPROVED
OAT
°C
52
45
40
35
30
25
20
0
-25
-40
BHT-407-FMS-25
407
0
137
139
140
140
140
140
140
140
140
137
407 AIRSPEED LIMITATIONS - KIAS
AIRSPEED LIMITATIONS – KNOTS – IAS
PRESSURE ALTITUDE FT x 1000
2
4
6
8
10
12
14
16
18
20
132
133
135
137
138
140
140
140
133
125
126
128
129
131
133
140
140
128
119
120
122
124
125
132
135
123
113
115
116
118
125
130
118
108
109
111
117
125
114
102
103
110
119
110
95
96
103
111
105
89
95
104
101
88
97
97
89
93
MAXIMUM AUTOROTATION VNE 100 KIAS
QUIET MODE VNE 100 KIAS
MAXIMUM QUIET MODE ALTITUDE IS 6000 FT HD
Airspeed limits shown are valid only for corresponding altitudes and temperatures.
Hatched areas indicate conditions which exceed approved temperature or density
altitude limitations.
407FS25-1-3
Figure 1-3. Placards and decals (typical)
17 DEC 2002
7
407-375-008-105
407-375-008-109
NP (POWER TURBINE RPM)
Quiet Cruise Mode
91.5%
Minimum
91.5 to 92.5%
Continuous operation
92.5%
Maximum continuous
Normal Operations
99%
Minimum
99 to 100%
Continuous operation
100%
Maximum continuous
NR (ROTOR RPM)
85%
Minimum (power off)
85 to 107%
Continuous operation (power off)
107%
Maximum (power off)
407_FMS_25_0001
Figure 1-4. Instrument Markings
8
Rev. 1
30 JUL 2008
TC APPROVED
BHT-407-FMS-25
Section 2
NORMAL PROCEDURES
2
2-4.
INTERIOR AND
PRESTART CHECK
QUIET NORMAL mode switch — NORMAL.
7.
QUIET NORMAL mode switch —
NORMAL.
8.
NR — 100% RPM.
9.
QUIET and ON annunciators —
Extinguished.
QUIET and ON annunciators illuminate and
extinguish with FADEC lights.
2-9.
2-6.
2-9-A.
2-6-E.
SYSTEMS CHECK
QUIET CRUISE MODE
CHECK
NOTE
If QUIET NORMAL mode switch is
cycled at less than 92% NR, increase
throttle to 100% NR to reset QUIET, ON
annunciator.
IN-FLIGHT OPERATIONS
IN QUIET CRUISE MODE
Flight at altitudes above 200 AGL and at
airspeeds above 50 KIAS:
QUIET NORMAL mode switch —
QUIET.
QUIET and ON annunciator —
Illuminated.
NR — 92%RPM.
1.
NR — 100% RPM.
2-10. DESCENT AND LANDING
2.
QUIET NORMAL mode switch —
QUIET.
2-10-B. IN QUIET CRUISE MODE
3.
QUIET and ON annunciators —
Illuminated.
4.
NR — 92% RPM.
5.
Throttle — Retard below 88% NR.
Confirm RPM warning illuminates
with audio.
QUIET NORMAL mode switch —
NORMAL. Monitor engine parameters.
Throttle — FLY detent position.
QUIET and ON annunciators —
Extinguished.
6.
Prior to descending below 200 AGL or 50
KIAS:
NR — 100% RPM
17 DEC 2002
9
BHT-407-FMS-25
TC APPROVED
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3-14. QUIET CRUISE MODE
OPERATION
3
NOTE
Landings into winds up to 35 knots
from azimuths of ± 45 degrees off nose
of helicopter have been demonstrated.
NOTE
In Quiet Cruise Mode, low rotor audio
and RPM caution light activated at 88%
NR.
If Quiet Cruise Mode fails engaged, plan
l a n d i n g i n t o w i n d . Tr a n s i e n t t o r q u e
excursions up to 100% during landing is
permitted.
Use of FADEC MAN mode will immediately
deselect Quiet Cruise Mode and reset low
rotor audio and RPM light activation point to
95% NR. To insure smooth transition to
FADEC MAN mode, match throttle position to
NG indication.
If either QUIET or ON segment is not
illuminated, return QUIET MODE switch to
NORMAL position.
Table 3-1. Warning (red) lights
PANEL
WORDING
FAULT CONDITION
CORRECTIVE ACTION
NR below 88%.
Reduce collective and ensure
throttle is in FLY detent
position. Light will extinguish
and audio will cease when NR
increases above 88%.
Warning (red) lights
RPM (with low RPM audio)
NR above 88%, QUIET light not Return to NORMAL mode.
illuminated.
Table 3-2. Caution (amber) and advisory (white/green) lights
PANEL
WORDING
FAULT CONDITION
HYDRAULIC SYSTEM
Hydraulic pressure below limit. Exit Quiet Cruise Mode,
returning to 100% NR. Verify
HYD SYS switch position.
Accomplish hydraulic system
failure procedure.
10
17 DEC 2002
CORRECTIVE ACTION
TC APPROVED
BHT-407-FMS-25
Section 4
PERFORMANCE
4
4-4.
HEIGHT – VELOCITY
ENVELOPE
The Height-velocity diagram (Figure 4-1)
defines conditions from which a safe landing
can be made on a smooth, level, firm surface,
following an engine failure. Limitations
respecting minimum airspeed and minimum
height above ground for Quiet Cruise Mode
operation are marked on the Height-velocity
diagram for clarity. The Height-velocity
diagram is valid only when helicopter gross
weight does not exceed limits of the Altitude
versus gross weight for height-velocity
diagram (Figure 4-2)
CAUTION
IF ENGINE FAILURE OCCURS DURING
FLIGHT CONDITIONS WITHIN HEIGHTVELOCITY DIAGRAM “AVOID AREA”,
SAFE LANDING MAY NOT BE
POSSIBLE.
4-7.
4-7-A.
CLIMB AND DESCENT
CLIMB
Reduce rate of climb data 100 feet per minute
when operating in Quiet Cruise Mode.
17 DEC 2002
11
BHT-407-FMS-25
TC APPROVED
HEIGHT-VELOCITY DIAGRAM
800
240
750
220
700
200
50 KIAS MINIMUM AIRSPEED
600
SKID HEIGHT - FEET
550
500
450
400
180
160
140
120
350
100
300
80
AVOID
250
200
SKID HEIGHT - METERS
650
200 FT MINIMUM HEIGHT ABOVE GROUND
150
60
40
100
20
50
0
0
0
10
20
30
40
50
60
70
80
90
100
110
120
INDICATED AIRSPEED - KNOTS
M407_FMS-25__FIG_4-1.WMF
Figure 4-1. Height – velocity diagram
12
17 DEC 2002
TC APPROVED
BHT-407-FMS-25
ALTITUDE VS GROSS WEIGHT
FOR HEIGHT-VELOCITY DIAGRAM
GROSS WEIGHT - KG x 100
13 14 15 16 17 18 19 20 21 22 23 24
5000 LB MAXIMUM INTERNAL GROSS WEIGHT
MAXIMUM OAT LIMIT
HEIGHT-VELOCITY
AVOID AREA APPLIES
Hp
- FT
-20
0
0
SE
AL
EV
EL
200
0
MINIMUM OAT LIMIT
400
0
600
0
800
0
10,0
00
6000 FT HD
-40
-20
0
20
OAT - °C
40
60
30
34
38
42
46
50
54
GROSS WEIGHT - LBS x 100
M407_FMS-25__FIG_4-2.WMF
Figure 4-2. Altitude vs gross weight for height – velocity diagram
17 DEC 2002
13/14
BHT-407-FMS-25
Section 5
WEIGHT AND BALANCE
5
5-1.
INTRODUCTION
This section presents loading information and
instructions necessary to ensure that flight
can be performed within approved gross
weight and center of gravity limitations as
defined in Section 1.
5-3.
GROSS WEIGHT
CENTER OF GRAVITY
5-3-B.
CENTER OF GRAVITY
Gross weight longitudinal center of gravity
and Gross weight lateral center of gravity
charts for Quiet Cruise operations are in
Limitations Section 1 .
For Quiet Cruise operations maintaining
longitudinal CG within limits can be achieved
by the following:
With helicopter Weight empty within
envelope (BHT-407-MM-2) the
helicopter will stay within Quiet Cruise
limits provided both pilot and co-pilot
seats are occupied. This assumes a
standard crew/passenger weight of 170
pounds (77.1 kilograms) and that fuel
and payload are adjusted to stay within
Maximum Gross Weight limit.
If co-pilot/forward passenger seat is
unoccupied then cabin payload must
be adjusted to maintain flight
envelope.
For helicopters operating without respect to
Weight empty envelope the pilot is
responsible for ensuring that when operating
in Quiet Cruise mode, helicopter weight and
CG are within limits.
For Quiet Cruise operations maintaining
Lateral CG within limits can be achieved by
the following:
Seats should be occupied such that
maximum asymmetric loading is no
more than one person (170 pounds
(77.1 kilograms)).
With this arrangement, a helicopter whose
basic lateral CG is ±0.3 inch (7.62 mm), will
remain within lateral limits.
17 DEC 2002
15/16
MANUFACTURER’S DATA
BHT-407-FMS-25
Section 1
SYSTEM DESCRIPTION
1
This kit incorporates a two position switch on
collective (Figure 1-1) perm itting pilot
selection of operation at 100% NR or 92% NR
(Quiet Cruise Mode). A two (2) segment
annunciator located on instrument panel
(Figure 1-2), when illuminated, displays
QUIET and ON. QUIET light is illuminated
when low rotor audio and RPM light activation
point is reset to 88% NR. ON light is
illuminated when FADEC is not in 100% NR
mode. Both segments are green and will
denote operation at 92% NR has been
selected. Dual tachometer has additional
markings to reflect permissible operation at
92% NP.
NOTE
Selection of FADEC switch to MAN
mode, for training purposes, while in
Quiet Cruise Mode, will immediately
deselect Quiet Cruise Mode and reset
low rotor audio and RPM light
activation point to 95% NR (triggering
low rotor audio and RPM light). QUIET
and ON segments will also extinguish.
Transition back to AUTO mode should
be accomplished at approximately
100% NP to reduce engine power
transients.
A FADEC system failure (FADEC FAIL
light with FADEC fail audio), while in
Quiet Cruise Mode, will retain
activation point of low rotor audio and
RPM light at 88%. QUIET segment will
remain illuminated and ON segment
will extinguish. Selection of FADEC
MAN mode will immediately deselect
Quiet Cruise Mode and reset low rotor
audio and RPM activation point to 95%
NR, extinguishing the QUIET segment.
1-55. NOISE LEVELS
1-55-A. FAR PART 36 STAGE 2
NOISE LEVEL
Flyover noise level in Quiet Cruise Mode for
the Model 407 is 81.3 dBA SEL.
1-55-B. CANADIAN
AIRWORTHINESS
MANUAL CHAPTER 516
AND ICAO ANNEX 16
NOISE LEVEL
Flyover noise level in Quiet Cruise Mode for
the Model 407 is 81.3 dBA SEL.
17 DEC 2002
17
BHT-407-FMS-25
MANUFACTURER’S DATA
Figure 1-1. Pilot collective stick
Figure 1-2. Quiet mode annunciator
18
17 DEC 2002
BHT-407-FMS-28
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
INCREASED INTERNAL GROSS WEIGHT
407-706-020
CERTIFIED
16 MARCH 1999
This supplement shall be attached to Model 407 Flight
Manual when INCREASED INTERNAL GROSS WEIGHT
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
2002
COPYRIGHT
BELL ® HELICOPTER TEXTRON
INC. AND BELL HELICOPTER
TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101
REISSUED — 16 DECEMBER 2002
BHT-407-FMS-28
NOTICE PAGE
PROPRIETARY RIGHTS NOTICE
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.
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
16 DEC 2002
BHT-407-FMS-28
LOG OF REVISIONS
Original ....................... 0................... 16 MAR 99
Reissue........................0 ................... 16 DEC 02
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
FLIGHT MANUAL
Title............................................................ 0
NP .............................................................. 0
A — B ........................................................ 0
C/D ........................................................... 0
1 — 92 ..................................................... 0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
16 DEC 2002
A
BHT-407-FMS-28
LOG OF TC APPROVED REVISIONS
Original ........................0 ...................16 MAR 99
B
16 DEC 2002
Reissue ....................... 0 ....................16 DEC 02
BHT-407-FMS-28
LOG OF FAA APPROVED REVISIONS
Original ....................... 0......................7 MAY 99
Reissue ....................... 0.................... 24 SEP 03
16 DEC 2002
C/D
TC APPROVED
BHT-407-FMS-28
Section 1
LIMITATIONS
1
1-6. WEIGHT AND CENTER
OF GRAVITY
For lateral CG limits, refer to Gross Weight
Lateral center of gravity limits chart
(Figure 1-2).
1-6-A. WEIGHT
1-7. AIRSPEED
Maximum approved internal gross weight for
takeoff and landing is 5250 pounds (2381
kilograms) or as shown in IGE hover
performance charts, Section 4.
V NE is 140 KIAS, sea level to 3000 feet H D .
Decrease V NE for ambient conditions in
accordance with AIRSPEED LIMITATIONS
Placards and decals (Figure 1-3).
CAUTION
LOADS THAT RESULT IN GW ABOVE
5,250 POUNDS (2381 KILOGRAMS)
SHALL BE CARRIED ON THE CARGO
HOOK.
1-6-B. CENTER OF GRAVITY
1-8. ALTITUDE
1-8-A. DENSITY
Maximum H D for takeoff, landing, and in
ground effect maneuvers is 11,000 feet (3353
meters).
1-8-B. DELETED
For longitudinal CG limits, refer to Gross
Weight Longitudinal center of gravity limits
charts (Figure 1-1).
16 DEC 2002
1
BHT-407-FMS-28
TC APPROVED
LONGITUDINAL C.G.
6200
120.5
6000
127.6
6000
5800
EXTERNAL LOAD ONLY
5600
5400
5250
5200
128.7
119.8
GROSS WEIGHT - POUNDS
5000
5000
4800
4600
4500
4400
4200
4000
3800
3600
3400
3200
3000
2800
2800
119.0
2600
2400
118
119
2650
120
121
122
123
124
125
126
127
128.0
129.0
128
129
130
FUSELAGE STATION - INCHES
M407_FMS-28__FIG_1-1_(1_OF_2).WMF
Figure 1-1. Gross Weight Longitudinal center of gravity limits (sheet 1 of 2)
2
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
LONGITUDINAL C.G.
2800
2700
3061
3241
2722
2600
EXTERNAL LOAD ONLY
2500
2400
2381
3042
2300
3268
GROSS WEIGHT - KILOGRAMS
2268
2200
2100
2041
2000
1900
1800
1700
1600
1500
1400
1300
1200
1100
3000
1270
3023
3025
1202
3050
3075
3100
3125
3150
3175
3200
3225
3251
3277
3250
3275
3300
FUSELAGE STATION - MILLIMETERS
M407_FMS-28__FIG_1-1_(2_OF_2).WMF
Figure 1-1. Gross Weight Longitudinal center of gravity limits (sheet 2 of 2)
16 DEC 2002
3
BHT-407-FMS-28
TC APPROVED
LATERAL C.G.
6200
-0.9
6000
1.4
6000
5800
5600
EXTERNAL LOAD ONLY
5400
MAX AIRSPEED 100 KIAS
5250
5200
-1.4
GROSS WEIGHT - POUNDS
5000
1.9
5000
MAX AIRSPEED
100 KIAS
4800
4600
MAX AIRSPEED
100 KIAS
-1.5
EXTERNAL
LOAD ONLY
2.0
EXTERNAL
LOAD ONLY
4400
4200
4000
3800
3600
3500
3500
3400
3200
3000
2800
2650
2600
-4.0
-2.5
3.0
4.0
3
4
2400
-5
-4
-3
-2
-1
0
1
2
5
BUTTOCK LINE - INCHES
M407_FMS-28__FIG_1-2_(1_OF_2).WMF
Figure 1-2. Gross Weight Lateral center of gravity limits (sheet 1 of 2)
4
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
LATERAL C.G.
2800
-23
36
2722
2700
2600
2500
EXTERNAL LOAD ONLY
MAX AIRSPEED 100 KIAS
2400
2381
GROSS WEIGHT - KILOGRAMS
2300
-35
2200
MAX AIRSPEED
100 KIAS
2100
EXTERNAL
LOAD ONLY
48
2268
MAX AIRSPEED
100 KIAS
52
-39
EXTERNAL
LOAD ONLY
2000
1900
1800
1700
1600
1588
1588
1500
1400
1300
1202
1200
-102
1100
-125
-100
-64
-75
-50
-25
0
25
50
76
102
75
100
125
BUTTOCK LINE - MILLIMETERS
M407_FMS-28__FIG_1-2_(2_OF_2).WMF
Figure 1-2. Gross Weight Lateral center of gravity limits (sheet 2 of 2)
16 DEC 2002
5
BHT-407-FMS-28
TC APPROVED
407 (5250 LB) AIRSPEED LIMITATIONS - KIAS
OAT
°C
52
45
40
35
30
25
20
0
-25
-40
PRESSURE ALTITUDE FT x 1000
0
137
139
140
140
140
140
140
140
140
137
2
4
6
8
10
12
14
16
18
20
132
133
135
137
138
140
140
140
133
123
125
128
129
131
133
140
140
128
113
116
118
121
124
132
135
123
104
106
109
112
123
130
118
99
100
102
111
125
114
93
94
101
114
110
86
87
94
102
105
80
86
95
101
79
88
93
80
86
MAXIMUM AUTOROTATION VNE 100 KIAS
Location: Adjacent to existing airspeed limitations placard (typical).
Airspeed limits shown are valid only for corresponding altitudes and temperatures.
Hatched areas indicate conditions which exceed approved temperature or density
altitude limitations.
407FS28-1-3
Figure 1-3. Placards and Decals (typical)
6
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
Section 2
NORMAL PROCEDURES
2
No change from basic manual.
16 DEC 2002
7
BHT-407-FMS-28
TC APPROVED
Section 3
EMERGENCY/MALFUNCTION PROCEDURES
3
No change from basic manual.
8
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
Section 4
PERFORMANCE
4
4-1. INTRODUCTION
Refer to appropriate performance charts in
accordance with optional equipment installed.
4-4. HEIGHT – VELOCITY
ENVELOPE
Altitude vs gross weight for height-velocity
diagram (Figure 4-1) and Height-velocity
(Figure 4-2) diagrams define conditions from
which a safe landing can be made on a
smooth, level, firm surface following an
engine failure. Height velocity diagram is valid
only when helicopter gross weight does not
exceed limits of the Altitude vs Gross Weight
diagram.
4-5. HOVER CEILING
NOTE
Hover performance charts are based on
100% ROTOR RPM.
Satisfactory stability and control have been
demonstrated in each area of the Hover
ceiling charts with winds as depicted on
Hover ceiling wind accountability chart (refer
to Basic Flight Manual).
Hover ceiling – in ground effect charts
(Figures 4-3 and 4-4) and Hover ceiling – out
of ground effect charts (Figures 4-5 and 4-6)
present hover performance as allowable
gross weight for conditions of H P and OAT.
These hovering weights are obtainable in zero
wind conditions. Each chart is divided into
two areas. Area A (non shaded area) of hover
ceiling charts presents hover performance
(relative to GW) for conditions where
adequate control margins exist for all relative
wind conditions up to 35 knots, for lateral CG
not exceeding ±2.5 inches (±63 mm); and up
to 17 knots, for lateral CG to ±4.0 inches (±102
mm); for hover, takeoff, and landing. Area B
(shaded area) of hover ceiling charts presents
hover performance (relative to GW) where
adequate control margins exist for relative
winds within ± 45° of nose of helicopter up to
35 knots, for lateral CG not exceeding ±2.5
inches (±63 mm); and up to 17 knots, for
lateral CG to ±4.0 inches (±102 mm); for hover,
takeoff, and landing.
4-7. CLIMB AND DESCENT
4-7-A. RATE OF CLIMB
Rate of climb (takeoff power) charts are
presented in Figure 4-7, and Rate of climb
(maximum continuous power) charts are
presented in Figure 4-8.
4-10. NOISE LEVELS
4-10-A. FAR PART 36 STAGE 2
NOISE LEVEL
Model 407 is certified as a Stage 2 helicopter
as prescribed in FAR Part 36, Subpart H, for
gross weights up to and including certificated
maximum takeoff and landing weight of 5250
pounds (2382 kilograms).
Certified flyover noise level for Model 407 is
85.5 dBA SEL.
16 DEC 2002
9
BHT-407-FMS-28
TC APPROVED
4-10-B. CANADIAN
AIRWORTHINESS MANUAL
CHAPTER 516 AND ICAO
ANNEX 16 NOISE LEVEL
by International Civil Aviation Organization
(ICAO) in Annex 16, Volume 1, Chapter 11, for
gross weights up to and including certified
maximum takeoff and landing weight of 5250
pounds (2382 kilograms).
Model 407 complies with noise emission
standards applicable to helicopter as set out
Flyover noise level for Model 407 is 85.5 dBA
SEL.
10
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
ALTITUDE VS GROSS WEIGHT
FOR HEIGHT-VELOCITY DIAGRAM
GROSS WEIGHT - KG x 100
13 14 15 16 17 18 19 20 21 22 23 24
14
,0
00
14,000 FT HD
00
D
S
OS
60
GR
C
40
G
EI
00
W
HT
LE
V
A
-2
0
00
SE
A
T
-F
Hp
S
EL
20
00
B
ON
GI
RE
-40
MAXIMUM INTERNAL GROSS WEIGHT = 5250 LB
80
00
10
,
MINIMUM OAT LIMIT
9000 FT HD
AT
UM O
00
0
12
MAXIM
,0
00
DEMONSTRATED
TO 9000 FEET
DENSITY ALTITUDE
-20
0
20
OAT - °C
40
60
30
34
38
42
46
50
54
GROSS WEIGHT - LBS x 100
M407_FMS-28__FIG_4-1.EMF
Figure 4-1. Altitude vs gross weight for height – velocity diagram
16 DEC 2002
11
BHT-407-FMS-28
TC APPROVED
HEIGHT-VELOCITY DIAGRAM
FOR GROSS WEIGHT REGIONS A TO D
800
240
E
LIN
LIN
E
IS
IS
E
LIN
TH
TH
W
LIN HIS
LO
IS
T
OW
BE
TH
W
EL
O
"B
EL
"A
"C "
200
AV
B
B"
400
350
140
120
100
300
SKID HEIGHT - METERS
450
E
SKID HEIGHT - FEET
160
ON
GI
RE
500
S
550
180
EA
AR
W
N"
LO
IO
OID
BE
G
RE
600
220
NOTE: LOW HOVER POINT IS
AT 6 FT SKID HEIGHT
"D "
N
GIO
650
RE
700
N
GIO
RE
750
80
250
200
60
AVOID
150
40
100
20
50
0
0
0
10
20
30
40
50
60
70
80
90
100
110
120
INDICATED AIRSPEED - KNOTS
M407_FMS-28__FIG_4-2.EMF
Figure 4-2. Height – velocity diagram
12
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
M407_FMS-28__FIG_4-3_(1_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 1 of 16)
16 DEC 2002
13
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-3_(2_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 2 of 16)
14
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
M407_FMS-28__FIG_4-3_(3_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 3 of 16)
16 DEC 2002
15
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-3_(4_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 4 of 16)
16
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(5_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 5 of 16)
16 DEC 2002
17
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(6_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 6 of 16)
18
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(7_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 7 of 16)
16 DEC 2002
19
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(8_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 8 of 16)
20
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-3_(9_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 9 of 16)
16 DEC 2002
21
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-3_(10_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 10 of 16)
22
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-3_(11_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 11 of 16)
16 DEC 2002
23
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-3_(12_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 12 of 16)
24
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(13_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 13 of 16)
16 DEC 2002
25
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(14_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 14 of 16)
26
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(15_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 15 of 16)
16 DEC 2002
27
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-3_(16_OF_16).EPS
Figure 4-3. Hover ceiling IGE – takeoff power (sheet 16 of 16)
28
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
M407_FMS-28__FIG_4-4_(1_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 1 of 16)
16 DEC 2002
29
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-4_(2_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 2 of 16)
30
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
M407_FMS-28__FIG_4-4_(3_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 3 of 16)
16 DEC 2002
31
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-4_(4_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 4 of 16)
32
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(5_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 5 of 16)
16 DEC 2002
33
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(6_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 6 of 16)
34
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(7_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 7 of 16)
16 DEC 2002
35
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
14
GROSS WEIGHT - 100 KG
18
20
22
24
26
60
00
0
-2
00
0
LE
VE
L
SE
A
Hp
-F
T
20
00
MAX GW
INTERNAL=5250 LB
40
00
0
20
OAT - °C
0 0
0 - 3 -2
-4
0
10
-20
40
60 32
36
40
44
48
52
GROSS WEIGHT - 100 LB
56
JETTISONABLE EXTERNAL=6000 LB
00
00
10
,0
80
C
-1
MINIMUM OAT
-°
AT
UM O
-40
AT
O
MAXIM
12
,
00
0
11,000 FT HD
16
60
M407_FMS-28__FIG_4-4_(8_OF_16).EMF
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 8 of 16)
36
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-4_(9_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 9 of 16)
16 DEC 2002
37
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-4_(10_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 10 of 16)
38
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-4_(11_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 11 of 16)
16 DEC 2002
39
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-4_(12_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 12 of 16)
40
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(13_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 13 of 16)
16 DEC 2002
41
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(14_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 14 of 16)
42
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(15_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 15 of 16)
16 DEC 2002
43
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
IN GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 4 FT (1.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS
M407_FMS-28__FIG_4-4_(16_OF_16).EPS
Figure 4-4. Hover ceiling IGE – maximum continuous power (sheet 16 of 16)
44
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
M407_FMS-28__FIG_4-5_(1_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 1 of 16)
16 DEC 2002
45
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-5_(2_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 2 of 16)
46
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
M407_FMS-28__FIG_4-5_(3_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 3 of 16)
16 DEC 2002
47
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-5_(4_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 4 of 16)
48
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(5_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 5 of 16)
16 DEC 2002
49
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(6_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 6 of 16)
50
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(7_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 7 of 16)
16 DEC 2002
51
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(8_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 8 of 16)
52
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-5_(9_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 9 of 16)
16 DEC 2002
53
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-5_(10_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 10 of 16)
54
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-5_(11_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 11 of 16)
16 DEC 2002
55
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-5_(12_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 12 of 16)
56
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(13_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 13 of 16)
16 DEC 2002
57
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(14_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 14 of 16)
58
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(15_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 15 of 16)
16 DEC 2002
59
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-5_(16_OF_16).EPS
Figure 4-5. Hover ceiling OGE – takeoff power (sheet 16 of 16)
60
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
M407_FMS-28__FIG_4-6_(1_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 1 of 16)
16 DEC 2002
61
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-6_(2_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 2 of 16)
62
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
M407_FMS-28__FIG_4-6_(3_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 3 of 16)
16 DEC 2002
63
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
M407_FMS-28__FIG_4-6_(4_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 4 of 16)
64
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(5_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 5 of 16)
16 DEC 2002
65
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(6_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 6 of 16)
66
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(7_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 7 of 16)
16 DEC 2002
67
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(8_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 8 of 16)
68
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-6_(9_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 9 of 16)
16 DEC 2002
69
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-6_(10_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 10 of 16)
70
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-6_(11_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 11 of 16)
16 DEC 2002
71
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
BASIC INLET
SNOW BAFFLES
M407_FMS-28__FIG_4-6_(12_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 12 of 16)
72
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE OFF
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(13_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 13 of 16)
16 DEC 2002
73
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(14_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 14 of 16)
74
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(15_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 15 of 16)
16 DEC 2002
75
BHT-407-FMS-28
TC APPROVED
HOVER CEILING
OUT OF GROUND EFFECT
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
SKID HEIGHT 40 FT (12.2 METER)
HEATER AND ANTI-ICE ON
PARTICLE SEPARATOR
SNOW BAFFLES
WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 50 LBS
M407_FMS-28__FIG_4-6_(16_OF_16).EPS
Figure 4-6. Hover ceiling OGE – maximum continuous power (sheet 16 of 16)
76
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
4000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
-4 0
14000
- 20
-3
12000
4
0
-10
OA
10000
T-
0
°C
8000
3
10
X
OA
20
T
2
LI
6000
MI
T
30
4000
PRESSURE ALTITUDE - 1000 METERS
5
MA
PRESSURE ALTITUDE - FEET
16000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(1_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 1 of 8)
16 DEC 2002
77
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
BASIC INLET
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
1000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
14000
-4
12000
OA
-3
T°C
-2
0
4
0
0
-1
0
10000
3
0
8000
10
20
6000
2
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(2_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 2 of 8)
78
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
PARTICLE SEPARATOR - PURGE OFF
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
3500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
-40
5
- 20
4
-1 0
12000
OA
T
10000
-°C
-3
0
0
3
10
8000
X
20
OA
2
T
MI
LI
6000
30
T
PRESSURE ALTITUDE - 1000 METERS
14000
MA
PRESSURE ALTITUDE - FEET
16000
4000
40
1
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(3_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 3 of 8)
16 DEC 2002
79
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER ON
PARTICLE SEPARATOR - PURGE OFF
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
14000
12000
OA
10000
-3
0
T°C
-2
-1
4
-4
0
0
3
0
0
8000
20
10
2
6000
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(4_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 4 of 8)
80
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
60 KIAS
HEATER OFF
BASIC INLET
SNOW DEFLECTOR
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
2500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
-4 0
14000
12000
5
-2 0
4
-1
M
AX
LI O
10000 MIT AT
OA
T-
-3
0
0
0
3
°C
10
8000
20
2
6000
30
4000
PRESSURE ALTITUDE - 1000 METERS
PRESSURE ALTITUDE - FEET
16000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(5_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 5 of 8)
16 DEC 2002
81
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
14000
4
12000
-4
OA
T
10000
-°C
-2
0
0
3
-3
0
-1
0
8000
0
2
10
6000
20
4000
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(6_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 6 of 8)
82
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
60 KIAS
HEATER OFF
PARTICLE SEPARATOR - PURGE OFF
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
2000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
6
18000
-40
14000
-2
0
12000
4
-1
M
AX
LI O
10000 MI AT
T
OA
0
-3
0
T°C
0
3
10
8000
20
2
6000
30
PRESSURE ALTITUDE - 1000 METERS
5
16000
PRESSURE ALTITUDE - FEET
10
4000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(7_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 7 of 8)
16 DEC 2002
83
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER ON
PARTICLE SEPARATOR - PURGE OFF
SNOW DEFLECTOR
TAKEOFF POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
20000
10
6
18000
PRESSURE ALTITUDE - FEET
14000
4
12000
-3
0
10000
OA
8000
-4
0
3
T°C
-2
0
-1
0
2
0
6000
10
4000
20
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-7_(8_OF_8).WMF
Figure 4-7. Rate of climb – takeoff power (sheet 8 of 8)
84
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
2500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
-30
14000
-2
-1
0
4
0
12000
0
OA
T-
10000
°C
3
10
8000
20
MA
2
X
6000
T
OA
PRESSURE ALTITUDE - FEET
-40
30
LI M
IT
4000
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(1_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 1 of 8)
16 DEC 2002
85
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
4
12000
-3
0
10000
OA
8000
-4
0
-2
0
T°C
3
-1
0
0
2
6000
10
20
4000
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(2_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 2 of 8)
86
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
60 KIAS
HEATER OFF
PARTICLE SEPARATOR - PURGE OFF
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
2000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
14000
-2
12000
0
-3
0
4
0
-1
0
OA
10000
T-
0
3
°C
10
8000
X
MA
6000
20
TL
OA
PRESSURE ALTITUDE - FEET
-4
30
IT
IM
4000
2
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
40
2000
50
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(3_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 3 of 8)
16 DEC 2002
87
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER ON
PARTICLE SEPARATOR - PURGE OFF
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
4
12000
-4
0
10000
3
-3
0
O
8000
-2
0
AT
-° C
-1
0
2
0
6000
10
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
20
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(4_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 4 of 8)
88
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
60 KIAS
HEATER OFF
BASIC INLET
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
1000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
-4
12000
-2
0
-1
10000
OA
T
0
4
-3
0
0
3
0
-°C
8000
10
2
6000
20
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
30
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(5_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 5 of 8)
16 DEC 2002
89
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER ON
BASIC INLET
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
4
12000
-4
0
-3
0
10000
-2
O
AT
-°C
3
0
-1
0
8000
0
2
6000
10
20
4000
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(6_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 6 of 8)
90
16 DEC 2002
TC APPROVED
BHT-407-FMS-28
RATE OF CLIMB
60 KIAS
HEATER OFF
PARTICLE SEPARATOR - PURGE OFF
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN ABOVE
500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
-4
0
4
-3
0
12000
-2
0
-1
0
10000
3
O
8000
0
AT
-° C
10
2
6000
20
4000
PRESSURE ALTITUDE - 1000 METERS
5
16000
1
30
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(7_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 7 of 8)
16 DEC 2002
91
BHT-407-FMS-28
TC APPROVED
RATE OF CLIMB
60 KIAS
HEATER ON
PARTICLE SEPARATOR - PURGE OFF
SNOW DEFLECTOR
MAXIMUM CONTINUOUS POWER
ENGINE RPM 100%
GENERATOR 180 AMPS
REDUCE RATE OF CLIMB 130 FT/MIN
FOR ANTI-ICE ON (5°C AND COLDER)
REDUCE RATE OF CLIMB 45 FT/MIN FOR PURGE ON
Gross Weight 5250 lb (2381 kg)
RATE OF CLIMB - M/SEC
0
1
2
3
4
5
6
7
8
9
10
20000
6
18000
PRESSURE ALTITUDE - FEET
14000
4
12000
-4
0
-3
0
10000
3
-2
0
O
8000
AT
-1
-°
C
0
2
0
6000
10
PRESSURE ALTITUDE - 1000 METERS
5
16000
4000
1
20
2000
0
0
200
400
600
800
1000
1200
1400
1600
1800
0
2000
RATE OF CLIMB - FT/MIN
M407_FMS-28__FIG_4-8_(8_OF_8).WMF
Figure 4-8. Rate of climb – maximum continuous power (sheet 8 of 8)
92
16 DEC 2002
BHT-407-FMS-31
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
INCREASED APC STARTER
GENERATOR LOAD
407-706-026
CERTIFIED
15 JUNE 2005
This supplement shall be attached to Model 407 Flight
Manual when Increased APC Starter Generator Load kit has
been installed.
Information contained herein supplements 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
COPYRIGHT
2005
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
15 JUNE 2005
BHT-407-FMS-31
NOTICE PAGE
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 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———15 JUN 2005
BHT-407-FMS-31
LOG OF REVISIONS
Original ..................... 0...................... 15 JUN 05
LOG OF PAGES
REVISION
NO.
PAGE
PAGE
REVISION
NO.
FLIGHT MANUAL
Title.............................................................0
NP ...............................................................0
A/B..............................................................0
C/D..............................................................0
E/F ..............................................................0
1 – 2 ............................................................0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
15 JUN 2005———A/B
BHT-407-FMS-31
LOG OF TC APPROVED REVISIONS
Original ..................... 0...................... 15 JUN 05
APPROVED
DATE
CHIEF, FLIGHT TEST
FOR
DIRECTOR — AIRCRAFT CERTIFICATION
TRANSPORT CANADA
15 JUN 2005———C/D
BHT-407-FMS-31
LOG OF FAA APPROVED REVISIONS
Original ..................... 0...................... 30 JUN 05
15 JUN 2005———E/F
TC APPROVED
BHT-407-FMS-31
Section 1
LIMITATIONS
1
1-12. ELECTRICAL
1-12-A. GENERATOR
1-20. INSTRUMENT MARKINGS
AND PLACARDS
Refer to Figure 1-5 for Instrument marking.
Continuous operation,
up to 10,000 feet HP
0 to 200 Amps
Maximum continuous up
to 10,000 feet HP
200 Amps
Continuous operation,
above 10,000 feet HP
0 to 180 Amps
Maximum continuous
above 10,000 feet HP
180 Amps
Transient, 2 minutes
200 to 300 Amps
Transient, 5 seconds
300 to 400 Amps
15 JUN 2005———1
BHT-407-FMS-31
TC APPROVED
* P/N 407-375-007-109
DC LOAD
0 to 200 Amps
Continuous operation
180 Amps
Maximum continuous above 10,000 FT Hp
200 Amps
Maximum
300 Amps
Maximum transient, 2 minutes
400 Amps
Maximum transient, 5 seconds
FUEL PRESSURE
8 PSI
Minimum
8 to 25 PSI
Continuous operation
25 PSI
Maximum
407FMS_31_0001
Figure 1-5. Instrument Marking
2———15 JUN 2005
BHT-407-FMS-CAA
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
UNITED KINGDOM REGISTERED
HELICOPTERS
CAA CERTIFIED
08 JANUARY 2002
This supplement shall be attached to the Bell Helicopter
Model 407 Flight Manual when the helicopter is registered in
the United Kingdom.
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
2002
COPYRIGHT
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell Helicopter
A Subsidiary of Textron Inc.
POST OFFICE BOX 482
FORT WORTH, TEXAS 76101
08 JANUARY 2002
BHT-407-FMS-CAA
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
08 JAN 2002
BHT-407-FMS-CAA
LOG OF REVISIONS
Original ....................0 ....................... 08 JAN 02
LOG OF PAGES
PAGE
REVISION
NO.
Title............................................................ 0
NP .............................................................. 0
PAGE
REVISION
NO.
A — B........................................................ 0
1 — 4 ......................................................... 0
5/6 ............................................................. 0
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
08 JAN 2002
A
BHT-407-FMS-CAA
CAA APPROVED
LOG OF APPROVED REVISIONS
Original .................... 0 ........................08 JAN 02
APPROVED:
CIVIL AVIATION AUTHORITY
SAFETY REGULATION GROUP
AVIATION HOUSE
SOUTH AREA
GATWICK AIRPORT
GATWICK
WEST SUSSEX RH6 OYR
B
08 JAN 2002
407-FMS-CAA.fm Page 1 Tuesday, January 4, 2005 4:40 PM
CAA APPROVED
BHT-407-FMS-CAA
Section 4
PERFORMANCE
4t
4-1. TAKEOFF DISTANCE
OVER 100-FOOT OBSTACLE
The Takeoff Distance Over 100-Foot Obstacle
c h a r t ( F i g u r e 4 - 1 ) p r o v i d e s ta k e o f f
performance data. The takeoff is initiated from
a stabilized 4-foot (1.2 meter) skid height
hover. Increase power smoothly to hover
power plus 20% torque or Takeoff Power,
whichever is less, and simultaneously start
nosedown pitch rotation so that the aircraft
accelerates along a flight path within the
takeoff corridor defined by the Height-Velocity
diagram (Figure 4-4 in Basic Manual). As the
helicopter goes through 50 KIAS, start nose
up rotation while increasing power to Takeoff
Power. With the aircraft starting to climb,
continue accelerating up to 65 KIAS. Engine
power limitations are imposed to preclude
unsafe nosedown attitude while in the flight
path required to remain clear of critical
he i ght -v e lo c it y l i mi ta ti o ns . Go od pi l ot
technique is required to achieve the
published takeoff performance. Wind factors
are not considered.
NOTE
Downwind takeoffs are not
recommended because the published
takeoff distance performance cannot
be achieved.
The power should be applied at a rate
sufficient to expedite the maneuver but
not so rapid as to overshoot the torque
value (approximately 6 seconds). Once
power is set, it should not be further
adjusted until the aircraft goes through
50 KIAS. At this airspeed, start nose up
rotation while increasing power to
Takeoff Power. While starting to climb,
continue accelerating up to 65 KIAS.
EXAMPLE:
What takeoff distance is required to clear a
1 0 0 - f o o t o b s ta c l e u n d e r t h e f o l l o w i n g
conditions:
OAT
HP
GW
20°C
1500 feet
4500 pounds
SOLUTION:
Enter the Takeoff Distance Over 100-Foot
Obstacle chart (Figure 4-1) at a pressure
altitude of 1500 feet, proceed horizontally to
t h e 2 0 ° C t e m p e r a t u r e l i n e . D r o p d ow n
vertically to the 4500 lb gross weight line and
move horizontally again to read a takeoff
distance of 1010 feet.
4-2.
PARTIAL POWER CLIMB
Torque limited partial power rate of climb
charts are presented for an aircraft with basic
inlet installed and with the heater and engine
anti-ice both OFF (Figure 4-2).
The recommended best rate of climb airspeed
is 60 KIAS.
08 JAN 2002
1
407-FMS-CAA.fm Page 2 Tuesday, January 4, 2005 4:40 PM
BHT-407-FMS-CAA
CAA APPROVED
EXAMPLE:
SOLUTION:
Find the maximum rate of climb that can be
attained using 65% torque under the following
conditions:
Enter the appropriate gross weight chart,
4000 lbs (Figure 4-2, Sheet 1 of 2). Starting at
a pressure altitude of 7000 feet, proceed
horizontally to the -10°C temperature line.
Drop down vertically and read a rate of climb
of 1225 feet per minute.
HEATER
ENGINE ANTI-ICE
OAT
HP
GW
2
08 JAN 2002
OFF
OFF
-10°C
7000 feet
4000 pounds
407-FMS-CAA.fm Page 3 Tuesday, January 4, 2005 4:40 PM
CAA APPROVED
BHT-407-FMS-CAA
TAKEOFF DISTANCE
OVER 100 FOOT OBSTACLE
CONDITIONS:
TECHNIQUE:
PRESSURE ALTITUDE - FEET
ROTOR RPM 100%
ROTATION AIRSPEED 50 KIAS
CLIMB AIRSPEED 65 KIAS
ZERO WIND
LEVEL ACCELERATION FROM 4 FT (1.2 M) SKID
HEIGHT USING HOVER POWER + 20% TORQUE
(NOT TO EXCEED 100%) UP TO 50 KIAS AND
CLIMB AT TAKEOFF POWER THEREAFTER
8000
6000
°C
T OA
-4 0
-30
-20
4000
-10
0
2000
10
20
30
40
50
0
-2000
AT
XO
MA
-4000
DISTANCE REQUIRED
TO CLEAR OBSTACLE - FEET
1300
1200
GROSS WEIGHT - LBS
5250
1100
5000
1000
4500
900
4000
800
700
3500 AND BELOW
Figure 4-1. Takeoff Distance Over 100-Foot Obstacle
08 JAN 2002
3
407-FMS-CAA.fm Page 4 Tuesday, January 4, 2005 4:40 PM
BHT-407-FMS-CAA
CAA APPROVED
PARTIAL POWER RATE OF CLIMB
RATE OF CLIMB LIMITED BY 65% TORQUE
60 KIAS
HEATER / ANTI-ICE OFF
BASIC INLET
ENGINE RPM 100%
GENERATOR 180 AMPS
GROSS WEIGHT = 4000 LBS
12000
11000
10000
8000
7000
-D
EG
C
6000
5000
-4
0
OA
T
-3
0
4000
-2
0
PRESSURE ALTITUDE - FEET
9000
10
0
-10
3000
30
40
20
2000
1000
0
750
800
850
900
950 1000 1050 1100 1150 1200 1250 1300 1350
RATE OF CLIMB - FT/MIN
Figure 4-2. Rate of Climb — Partial Power (Sheet 1 of 2)
4
08 JAN 2002
407-FMS-CAA.fm Page 5 Tuesday, January 4, 2005 4:40 PM
CAA APPROVED
BHT-407-FMS-CAA
PARTIAL POWER RATE OF CLIMB
RATE OF CLIMB LIMITED BY 65% TORQUE
60 KIAS
HEATER / ANTI-ICE OFF
BASIC INLET
ENGINE RPM 100%
GENERATOR 180 AMPS
GROSS WEIGHT = 5000 LBS
12000
-4 0
0
10
-3 0
-2 0
0
-1
11000
10000
8000
20
7000
6000
30
5000
40
4000
3000
O
T
A
EG
-D
C
-4
0
2000
1000
0
450
475
500
525
550
575
600
625
0
10
-2
0
0
-3
-1
0
PRESSURE ALTITUDE - FEET
9000
650
675
700
725
750
RATE OF CLIMB - FT/MIN
Figure 4-2. Rate of Climb — Partial Power (Sheet 2 of 2)
08 JAN 2002
5/6
BHT-407-FMS-IAC AR
3q/4fl7
ROTORCRAFT
FLIGHT MANUAL
SUPPLEMENT
INTERSTATE AVIATION COM M ITTEE —
AVIATION REGISTER
COMMONWEALTH OF INDEPENDENT
STATES
407-706-021
CERTIFIED
This supplement shall be attached to Model 407 Flight
Manual when IAC-AR kit is installed and helicopter is
registered in the Commonwealth of Independent States.
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
1999
COPYRIGHT
BELL® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON INC.
A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
Bell H
TEXTRON
A Subsidiary of Textron Inc.
POST OFFICE BOX 482 • FORT WORTH, TEXAS 78101
20 MAY 1999
BHT-407-FMS-IAC AR
NOTICE PAGE
PROPRIETARY RIGHTS NOTICE
Manufacturer's Data portion of this manual 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.
Add itional 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-407FMS-IAC AR
LOG OF REVISIONS
Original
0
.20 MAY 99
LOG OF PAGES
PAGE
REVISION
NO.
Title
0
NP
A—B
0
i/u
0
0
0
0
1—6
7/8
9/10
PAGE
REVISION
NO.
NOTE
Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of
superseded pages.
A
BHT-407-FMS-IAC AR
DOT APPROVED
LOG OF DOT APPROVED REVISIONS
Original
0
20MAY99
APPROVED
HIEF, FLIGH
FOR
DIRECTOR AIRCRAFT CERTIFICATION BRANCH
DEPARTMENT OF TRANSPORT
B
DATE
/2
DOT APPROVED
BHT-407-FMS-IAC AR
GENERAL INFORMATION
Interstate Aviation Committee — Aviation Register, Commonwealth of Independent States kit
consists of a metric altimeter for flight reference in metrically measured airspace environments,
provisions for crew and passenger oxygen system and associated equipment for use, a VHF
emergency radio communications system (located in bracketed compartment on interior side of
pilot door), and a fire detection system with aural and visual annunciators.
i/u
DOT APPROVED
BHT-407-FMS-IAC AR
Section 1
LIMITATIONS
1-1.
INTRODUCTION
Refer to Manufacturer's Data (BHT-4O7MD-1)
for metric equivalency conversions of altitude
limitations not covered in this supplement.
1-8.
ALTITUDE
1-17. FUEL AND OIL
CIS fuels which meet CIS specification
GOST#10227 (TS1 and RT) are suitable for
use with this helicopter.
CIS fuel TS-1 may be used at all ambient
temperatures.
NOTE
CIS fuel RT usage is limited to ambient
Approved oxygen equipment must be
temperatures of -32° C (-25° F) or above.
operations above 2400 meters H with
passengers, or 3000 meters H for crew
only.
Anti-icing fuel additives are required for all
fuels when ambient temperatures are below
installed on helicopter prior to
50 C (400 F).
NOTE
1-10. ATCROUTING
Additives Fluid I (Ethylene Glycol
Helicopter flights within CIS airspace are
allowed only along routes covered by ATC
ground facilities using RBS mode.
Permits to fly along routes covered by ATC
ground facilities using UVD (Russian) mode
monoethyl ether = GOST 8313) and
Fluid IM (a mix or 50% Fluid I and 50%
methyl alcohol = TU-6-10-1458) with a
concentration of 0.1 — 0.3% by volume
may be used for anti-icing with above
fuels.
shall be obtained from ATC service.
1-20.
Refer
INSTRUMENT MARKINGS
AND PLACARDS
to Figure
1-3 for Placards
Figure 1-3
Refer to Figure
Figure 1-5
1-5
and decals.
for Instrument markings.
1
BHT-407-FMS-IAC AR
DOT APPROVED
FOR OPERATING TEMPERATURES BELOW 5°C140 F° OAT,
C.I.S. RT AND TS-1 (GOST 10277)FUEL MUST CONTAIN
FLUID I (GOST 8313) OR FLUID IM (TU-6-1 0-14458) ADDITIVE.
CONCENTRATION TO BE 0.1 - 0.3% BY VOLUME.
Typical
Figure 1-3. Placards and decals
2
DOT APPROVED
BHT-407-FMS-IAC AR
4O7IACAR-1-5
Typical
Figure 1-5. Instrument markings
3
BHT-407-FMS-IAC AR
DOT APPROVED
Section 2
NORMAL PROCEDURES
2-1 .
INTRODUCTION
Refer to Manufacturer's Data (BHT-407-MD-1)
for metric equivalency conversions not
covered in this supplement.
2-4.
INTERIOR AND PRESTART
CHECK
Oxygen system and associated
IN-FLIGHT OPERATIONS
2-9.
Use metric altimeter as reference instrument
when helicopter is operated in metric air
space.
ENG ANTI ICE and PITOT HEATER
switches — ON, in visible moisture when
ambient temperature is at or below 5° C
(400 F).
equipment — Check.
FIRE DET TEST switch — Press.
Verify following:
ENGINE FIRE light illuminates (with
audio).
NOTE
At altitudes above 2400 meters H with
passengers, or 3000 meters H with
crew only, use oxygen masks.
2-9-A.
ICING CONDITIONS
FIRE DET TEST switch — Release.
Verify following:
ENGINE FIRE light extinguishes (audio
off).
2-5.
ENGINE START
Helicopter must exit any inadvertent entry into
icing conditions as soon as practical.
2-9-B. MANEUVERS
Do not exceed 600 per second Yaw rate
(rotation) in hover.
ENG ANTI ICE switch — ON.
PITOT HEATER switch — ON.
In powered flight, do not exceed sink rate
(vertical speed) of 600 feet per minute (3
AMPS indicator movement — Check.
meters per second) at airspeeds of 25 KIAS or
below.
Do not exceed bank angles of 35° with
passengers, or 60° with crew only.
2-11. ENGINE SHUTDOWN
In remote areas, ensure battery is being
recharged prior to engine shutdown by
increasing electrical load (i.e. using landing
light), and monitoring AMPS indicator. This
will reduce possibility of having an uncharged
battery and subsequent failure to start engine.
4
BHT-407-FMS-IAC AR
DOT APPROVED
213. COLD WEATHER
OPERATIONS
Cover engine air intake and exhaust ports
when parked during snow (or expected snow)
conditions.
Remove all snow and/or ice in vicinity of
engine air intake and exhaust ports prior to
If second attempt at engine start is required
(due to reaching starting limits during first
attempt) wait 3 minutes prior to next start
attempt to allow battery heat to distribute
through battery and engine heat to flow into
lubrication system.
Prior to embarking passengers, ground run
helicopter to warm cabin interior and
minimize fogging of windows.
removing covers and initiating engine start.
WARNING
CAUTION
PRIOR TO TAKEOFF, ENSURE
BLADE AND WINDSHIELD COVERS
MAY FREEZE IN PLACE AND DAMAGE
MAY RESULT DURING REMOVAL
LANDING SKIDS ARE NOT FROZEN TO
GROUND.
PROCESS.
CAUTION
NOTE
attention should be paid
ABRUPT CONTROL INPUT CHANGES
during preflight checks to those areas
subject to ice accumulation.
EXCESSIVELY WHEN PARKED ON ICE.
Particular
Engine starts are improved by keeping battery
in a warm environment prior to use.
NOTE
If possible, use an external power
CAN CAUSE HELICOPTER TO YAW
If operating in remote areas for extended
periods, perform periodic engine restarts to
maintain gearbox oil temperatures and battery
charge.
Of extensive operations at temperatures near
source for first start of day.
.4Qo C
One fuel boost pump can be used
using MIL-L-7808 lubricating oil in engine.
during start to conserve battery power.
Normal transmission and engine oil pressure
(4O0 F) are expected and starting is
difficult, consideration should be given to
Reference Allison Operation and Maintenance
Manual for procedures related to changing oil
types.
may be exceeded during start. Stabilize
engine at idle until minimum temperature and
pressure limits are reached.
5
BHT-407-FMS-IAC AR
DOT APPROVED
Section 3
EMERGENC V/MALFUNCTION PROCEDURES
3-4.
ENGINE FIRE DETECTION
Refer to Table 3-1 for wording, conditions, and
associated corrective action for engine fire
detection system.
Table 3-1. Warning (red) lights
PANEL
WORDING
ENGINE FIRE
(with audio)
FAULT CONDITION
CORRECTIVE ACTION
Excessive temperature
condition in engine
compartment
Immediately enter autorotation.
Throttle — Close.
FUEL VALVE switch — OFF.
If time permits, FUEL BOOST/XFR
circuit breaker switches — OFF.
Execute a normal autorotation and
landing
BATT switch — OFF.
NOTE
Do not restart engine until cause of fire has been determined and corrected.
Section 4
PERFORMANCE
No change from basic manual.
6
MANUFACTURER'S DATA
BHT-407-FMS-IAC AR
Section 5
WEIGHT AND BALANCE
No change from basic manual.
7/8
MANUFACTURER'S DATA
BHT-407-FMS-IAC AR
Section 1
SYSTEM DESCRIPTION
1-17. INSTRUMENT OPERATION
1-17-D. FLIGHT DATA RECORDER
Event Signals:
L Engine out
2.
Generator failure
Operations with commercial Flight Data
3. Low rotor RPM
parameters:
4. Transmission oil pressure
Recorder (FDR) must use following
Analog:
1.
Barometric altitude (from metric
altimeter)
2. Heading
3.
Vertical G-load
5. Transmission oil temperature
6. Engine fire
7. Radio ON.
1-48. EMERGENCY
COMMUNICATION SYSTEM
4. DC voltage
5. Time
148A. VHF EMERGENCY RADIO
6.
Engine torque
VHF emergency radio and its instructions for
7.
Engine RPM.
use are located in pilot access door
compartment.
9/10
BHT-407-MD-1
ROTORCRAFT
MANUFACTURER’S DATA
BHT-407-MD-1
ROTORCRAFT
MANUFACTURER’S DATA
COPYRIGHT NOTICE
COPYRIGHT
2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
REISSUE — 9 DECEMBER 2002
REVISION 4 — 30 APRIL 2008
BHT-407-MD-1
NOTICE PAGE
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 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. 2—31 JAN 2007
BHT-407-MD-1
LOG OF REVISIONS
Original ......................0 ......................20 SEP 96
Revision.....................1 ..................... 14 APR 97
Revision.....................2 ......................24 JUN 97
Revision.....................3 ..................... 05 AUG 98
Revision.....................4 ..................... 20 DEC 99
Revision.....................5 ......................15 FEB 00
Reissue ..................... 0 ......................09 DEC 02
Revision .................... 1 ...................... 29 JUN 05
Revision .................... 2 ...................... 31 JAN 07
Revision .................... 3 ..................... 26 MAR 07
Revision .................... 4 ......................30 APR 08
LOG OF PAGES
PAGE
REVISION
NO.
Cover............................................................ 2
Title .............................................................. 4
NP................................................................. 2
A/B................................................................ 4
i/ii.................................................................. 2
1-1 – 1-13 ..................................................... 2
1-14............................................................... 4
1-15 – 1-19 ................................................... 2
1-20 – 1-23 ................................................... 4
1-24............................................................... 2
1-25............................................................... 4
1-26............................................................... 2
1-27 – 1-28 ................................................... 3
1-29 – 1-68 ................................................... 2
PAGE
REVISION
NO.
1-69 .............................................................. 4
1-70 – 1-76 ................................................... 2
1-77 – 1-78 ................................................... 4
1-79 – 1-106 ................................................. 2
1-107/1-108 .................................................. 2
2-1 – 2-6 ....................................................... 2
2-7 – 2-16 ..................................................... 4
2-17/2-18 (Deleted) ..................................... 4
3-1/3-2 .......................................................... 0
3-3 – 3-8 ....................................................... 0
3-9/3-10 ........................................................ 0
4-1/4-2 .......................................................... 0
4-3 – 4-28 ..................................................... 0
NOTE
Revised text is indicated by a black vertical line. A revised page with only a vertical line next to
the page number indicates that text has shifted or that non-technical correction(s) were made
on that page. Insert latest revision pages; dispose of superseded pages.
30 APR 2008—Rev. 4———A/B
BHT-407-MD-1
GENERAL INFORMATION
The Manufacturer’s Data is provided for use in conjunction with the basic Flight Manual and
optional equipment supplements, as applicable. This manual contains useful information to
familiarize the operator with the helicopter and its systems, to facilitate ground handling and
servicing procedures, and to assist in flight planning and operations.
The Manufacturer’s Data is divided into four sections as follows:
Section 1 — SYSTEMS DESCRIPTION
Section 2 — HANDLING AND SERVICING
Section 3 — CONVERSION CHARTS AND TABLES
Section 4 — EXPANDED PERFORMANCE
31 JAN 2007—Rev. 2———i/ii
MANUFACTURER’S DATA
BHT-407-MD-1
Section 1
SYSTEM DESCRIPTION
1
TABLE OF CONTENTS
Subject
Paragraph
Number
Page
Number
Introduction ............................................................................................
Helicopter Description...........................................................................
Principal Dimensions ............................................................................
Location References..............................................................................
Fuselage Stations ..............................................................................
Water Lines.........................................................................................
Buttock Lines .....................................................................................
General Arrangement ............................................................................
Crew Compartment............................................................................
Passenger Compartment ..................................................................
Baggage Compartment .....................................................................
Tailboom .............................................................................................
Instrument Panel, Console, and Pedestal ...........................................
Instrument Panel................................................................................
Overhead Console and Pedestal ......................................................
Engine Instruments ...............................................................................
Instrument Operation ........................................................................
Power-on BIT ..................................................................................
Commanded BIT ............................................................................
Exceedance Monitoring.................................................................
Check Instrument (CHECK INSTR) Light.............................................
Clock .......................................................................................................
Clock Operation .................................................................................
Universal Time Setting ..................................................................
Local Time Setting .........................................................................
Elapsed Time Count Up.................................................................
Elapsed Time Count Down............................................................
Flight Time Reset ...........................................................................
Flight Time Alarm Set ....................................................................
Flight Time Alarm Display .............................................................
Test Mode .......................................................................................
Caution and Warning System ...............................................................
Power Plant ............................................................................................
Engine Controls — FADEC System .....................................................
FADEC System...................................................................................
FADEC System — Operation ............................................................
Power Up Mode and Built-in-test......................................................
1-1 ...........
1-2 ...........
1-3 ...........
1-4 ...........
1-4-A .......
1-4-B .......
1-4-C .......
1-5 ...........
1-5-A .......
1-5-B .......
1-5-C .......
1-5-D .......
1-6 ...........
1-6-A .......
1-6-B .......
1-7 ...........
1-7-A .......
1-7-A-1 ....
1-7-A-2 ....
1-7-A-3 ....
1-8 ...........
1-9 ...........
1-9-A .......
1-9-A-1 ....
1-9-A-2 ....
1-9-A-3 ....
1-9-A-4 ....
1-9-A-5 ....
1-9-A-6 ....
1-9-A-7 ....
1-9-A-8 ....
1-10 .........
1-11 .........
1-12 .........
1-12-A .....
1-12-B .....
1-12-C .....
1-7
1-7
1-7
1-7
1-7
1-7
1-11
1-11
1-13
1-14
1-14
1-14
1-15
1-15
1-15
1-20
1-20
1-20
1-20
1-20
1-21
1-21
1-22
1-22
1-22
1-22
1-22
1-23
1-23
1-23
1-23
1-23
1-25
1-25
1-25
1-27
1-27
31 JAN 2007—Rev. 2———1-1
BHT-407-MD-1
MANUFACTURER’S DATA
TABLE OF CONTENTS (CONT)
Subject
Start in Auto Mode.............................................................................
Alternate Start — AUTO MODE ........................................................
Start in Manual Mode.........................................................................
In-flight Auto Mode Operation ..........................................................
FADEC System Faults .......................................................................
Category 1 — FADEC Fail/FADEC Manual — Direct Reversion
to Manual System (DRTM) ................................................................
The FADEC System Will Fail to the Manual Mode as Follows:..
Category 2 — FADEC Degraded.......................................................
Category 3 — FADEC Fault...............................................................
Category 4 — Restart Fault...............................................................
FADEC Reversionary Governor (FADEC Software Version 5.356)
Engine Overspeed Protection ..........................................................
Engine Shutdown ..............................................................................
HMU Manual Piston Parking Procedure ..........................................
Category 5 — Maintenance Advisory, FADEC System Faults —
Engine Shutdown ..............................................................................
Checking FADEC Fault Codes..........................................................
Engine Run Fault Codes — Procedure For Viewing...................
FADEC Fault Codes — Procedure to Determine Last Engine
Run Faults From Current Faults...................................................
FauLt Code Charts — Use of ........................................................
Clearing FADEC Fault Codes ...........................................................
Current Faults ................................................................................
Last Engine Run Faults/Exceedances .........................................
Accumulated Faults/ Exceedances ..............................................
FADEC Training in Manual Mode .....................................................
Cockpit Procedural Training on Ground (Engine Not Running) ...
Flight Training in Manual Mode........................................................
Simulated FADEC Failure Training (In-flight)..................................
Combined Engine Filter Assembly (CEFA) .........................................
Power Plant Ignition System ................................................................
Power Plant Temperature Measurement System ...............................
Power Plant Compressor Bleed Air System .......................................
Engine Oil System .................................................................................
Engine Auto Relight ..............................................................................
Auto Relight Caution Light ...................................................................
Start in AUTO Mode...........................................................................
Engine Out in Auto Mode..................................................................
Start and Continuous Operation in Manual Mode ..........................
Engine ANTI ICE Switch........................................................................
Engine Indicators...................................................................................
Torque Gauge ....................................................................................
Torque Gauge — Range Marking .................................................
Engine Torque Exceedance ..........................................................
1-2———Rev. 2—31 JAN 2007
Paragraph
Number
Page
Number
1-12-D .....
1-12-E .....
1-12-F .....
1-12-G.....
1-12-H .....
1-27
1-31
1-31
1-32
1-33
1-12-I.......
1-12-I-1 ...
1-12-J......
1-12-K .....
1-12-L .....
1-12-M.....
1-12-N .....
1-12-O.....
1-12-P .....
1-34
1-34
1-43
1-43
1-43
1-44
1-44
1-46
1-46
1-12-Q.....
1-12-R .....
1-12-R-1..
1-47
1-47
1-47
1-12-R-2..
1-12-R-3..
1-12-S .....
1-12-S-1 ..
1-12-S-2 ..
1-12-S-3 ..
1-12-T .....
1-12-U .....
1-12-V .....
1-12-W ....
1-13.........
1-14.........
1-15.........
1-16.........
1-17.........
1-18.........
1-19.........
1-19-A .....
1-19-B .....
1-19-C .....
1-20.........
1-21.........
1-21-A .....
1-21-A-1..
1-21-A-2..
1-48
1-48
1-55
1-55
1-56
1-56
1-56
1-56
1-56
1-57
1-57
1-60
1-60
1-60
1-61
1-61
1-63
1-63
1-63
1-64
1-64
1-64
1-64
1-64
1-65
MANUFACTURER’S DATA
BHT-407-MD-1
TABLE OF CONTENTS (CONT)
Subject
Paragraph
Number
Page
Number
Gas Producer Gauge (NG) .................................................................
Dual Tach Gauge................................................................................
Measured Gas Temperature Gauge .................................................
Engine Oil Temperature/Pressure Gauge ........................................
Engine Out Warning Light and Horn....................................................
FADEC in Auto Mode .........................................................................
FADEC in Manual Mode.....................................................................
Engine Overspeed Warning Light ........................................................
Engine Chip Caution Light....................................................................
Fuel System Description .......................................................................
Fuel System Operation......................................................................
Left Fuel Boost/XFR Alternate Electrical Circuit.............................
Fuel System Controls ........................................................................
Fuel Valve Switch...............................................................................
Left and Right Fuel Boost/XFR Switch ............................................
Fuel Cell Drain Switches ...................................................................
Fuel System Indicators......................................................................
Fuel System Capacity........................................................................
Fuel Quantity Gauging System (FQGS) ...........................................
Fuel Quantity Signal Conditioner .....................................................
Signal Conditioner Built-in-test (BIT) ...............................................
Power-up BIT ..................................................................................
Continuous BIT ..............................................................................
Fuel Quantity Calculation..................................................................
Fuel Quantity Gauge..........................................................................
Fuel Quantity Button .........................................................................
Fuel Pressure/Ammeter Gauge ........................................................
Fuel Valve Light .................................................................................
L/Fuel and R/Fuel Boost Lights ....................................................
L/Fuel XFR and R/Fuel XFR Lights ...............................................
Fuel Filter Light ..................................................................................
Fuel Low Light....................................................................................
Retirement Index Number (RIN)............................................................
Transmission..........................................................................................
Freewheel Assembly .........................................................................
Transmission Oil System ..................................................................
Transmission Indicators ...................................................................
Transmission Oil Temperature and Pressure Gauge .....................
Transmission Oil Temperature Gauge .........................................
Transmission Oil Pressure Light..................................................
Transmission Chip Annunciator ......................................................
Rotor System..........................................................................................
Main Rotor Hub and Blades ..............................................................
Tail Rotor Hub and Blades ................................................................
Tail Rotor Gearbox.............................................................................
1-21-B .....
1-21-C .....
1-21-D .....
1-21-E .....
1-22 .........
1-22-A .....
1-22-B .....
1-23 .........
1-24 .........
1-25 .........
1-25-A .....
1-25-B .....
1-25-C .....
1-25-D .....
1-25-E .....
1-25-F......
1-25-G .....
1-25-H .....
1-25-I.......
1-25-J......
1-25-K .....
1-25-K-1 ..
1-25-K-2 ..
1-25-L......
1-25-M.....
1-25-N .....
1-25-O .....
1-25-P .....
1-25-P-1 ..
1-25-P-2 ..
1-25-Q .....
1-25-R .....
1-26 .........
1-27 .........
1-27-A .....
1-27-B .....
1-27-C .....
1-27-D .....
1-27-D-1 ..
1-27-D-2 ..
1-27-E .....
1-28 .........
1-28-A .....
1-28-B .....
1-28-C .....
1-66
1-66
1-66
1-68
1-68
1-68
1-68
1-68
1-69
1-69
1-69
1-73
1-73
1-73
1-73
1-74
1-74
1-74
1-74
1-74
1-76
1-76
1-76
1-76
1-76
1-77
1-77
1-77
1-77
1-77
1-78
1-79
1-79
1-79
1-79
1-81
1-81
1-83
1-83
1-83
1-83
1-83
1-83
1-83
1-86
31 JAN 2007—Rev. 2———1-3
BHT-407-MD-1
MANUFACTURER’S DATA
TABLE OF CONTENTS (CONT)
Subject
Paragraph
Number
Page
Number
Tail Rotor Chip Light .........................................................................
Rotor System Indicators .......................................................................
Dual Tach Gauge ...............................................................................
RPM Light and Warning Horn ...........................................................
NR Less than 95% ..........................................................................
NR 107% or Greater........................................................................
Flight Control System ...........................................................................
Rotor Controls ...................................................................................
Main Rotor ..........................................................................................
Cyclic ..............................................................................................
Collective ........................................................................................
Tail Rotor ............................................................................................
Airspeed Actuated Pedal Stop System............................................
Hydraulic System...................................................................................
Hydraulic Indicators ..........................................................................
Hydraulic Filter Indicators ................................................................
Hydraulic System Light.....................................................................
Electrical System ...................................................................................
External Power...................................................................................
Battery Switch....................................................................................
Battery Charging (17 and 28 Amp/Hour Batteries) .........................
Generator Switch ...............................................................................
Start Switch ........................................................................................
Electrical System Indicators.............................................................
Fuel Pressure/DC Ammeter ..............................................................
Ammeter — 200 Amps Max Scale ................................................
Ammeter — 400 Amps Max Scale ................................................
Voltmeter ............................................................................................
Battery Hot Annunciator ...................................................................
Battery Relay Annunciator................................................................
Start Annunciator ..............................................................................
Gen Fail Annunciator ........................................................................
Pitot Static System ................................................................................
Basic Flight Instruments.......................................................................
Airspeed Indicator .............................................................................
Altimeter .............................................................................................
Vertical Speed Indicator (VSI)...........................................................
Inclinometer .......................................................................................
Optional Flight Instruments..............................................................
Attitude Indicator ...............................................................................
Directional Gyro.................................................................................
Turn-and-Slip Indicator .....................................................................
Encoding Altimeter............................................................................
Navigation Systems and Instruments..................................................
Magnetic Compass ............................................................................
1-28-D .....
1-29.........
1-29-A .....
1-29-B .....
1-29-B-1..
1-29-B-2..
1-30.........
1-30-A .....
1-30-B .....
1-30-B-1..
1-30-B-2..
1-30-C .....
1-30-D .....
1-31.........
1-31-A .....
1-31-B .....
1-31-C .....
1-32.........
1-32-A .....
1-32-B .....
1-32-C .....
1-32-D .....
1-32-E .....
1-32-F .....
1-32-G.....
1-32-G-1..
1-32-G-2..
1-32-H .....
1-32-I.......
1-32-J......
1-32-K .....
1-32-L .....
1-33.........
1-34.........
1-34-A .....
1-34-B .....
1-34-C .....
1-34-D .....
1-34-E .....
1-34-F .....
1-34-G.....
1-34-H .....
1-34-I.......
1-35.........
1-35-A .....
1-86
1-86
1-86
1-86
1-86
1-86
1-87
1-87
1-87
1-87
1-87
1-89
1-89
1-89
1-90
1-90
1-90
1-90
1-94
1-94
1-95
1-95
1-96
1-96
1-96
1-96
1-96
1-96
1-97
1-97
1-97
1-97
1-97
1-97
1-97
1-99
1-99
1-99
1-99
1-99
1-99
1-99
1-99
1-99
1-99
1-4———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
TABLE OF CONTENTS (CONT)
Subject
Paragraph
Number
Page
Number
Optional Navigation Equipment .......................................................
Horizontal Situation Indicator (HSI)..............................................
Global Positioning System (GPS) ................................................
VOR Indicator .................................................................................
Automatic Direction Finder (ADF) ................................................
ATC Transponder...........................................................................
VHF NAV/COMM System ...............................................................
Three or Five Place Intercommunication System...............................
Normal Mode ......................................................................................
Isolate Mode .......................................................................................
Private Mode.......................................................................................
Aft Cabin Mode...................................................................................
Emergency Communication System....................................................
Avionics Master Switch.........................................................................
Miscellaneous Instruments...................................................................
Outside Air Temperature Indicator...................................................
Hourmeter...........................................................................................
Ventilation and Defog System ..........................................................
Lighting System .................................................................................
Cockpit Utility Light ...........................................................................
Instrument Panel and Associated Lighting .....................................
Aft Cabin Lighting ..............................................................................
Landing Lights ...................................................................................
Position Lights ...................................................................................
Anticollision Light..............................................................................
Emergency Equipment ..........................................................................
Portable Fire Extinguisher ....................................................................
First Aid Kit.............................................................................................
Pointer 4000 ELT....................................................................................
1-35-B .....
1-35-B-1 ..
1-35-B-2 ..
1-35-B-3 ..
1-35-B-4 ..
1-35-B-5 ..
1-35-B-6 ..
1-36 .........
1-36-A .....
1-36-B .....
1-36-C .....
1-36-D .....
1-37 .........
1-38 .........
1-39 .........
1-39-A .....
1-39-B .....
1-39-C .....
1-39-D .....
1-39-E .....
1-39-F......
1-39-G .....
1-39-H .....
1-39-I.......
1-39-J......
1-40 .........
1-41 .........
1-42 .........
1-43 .........
1-100
1-100
1-100
1-100
1-101
1-101
1-101
1-102
1-103
1-103
1-103
1-104
1-104
1-104
1-104
1-104
1-104
1-104
1-105
1-105
1-105
1-105
1-105
1-106
1-106
1-106
1-106
1-106
1-106
FIGURES
Subject
Principal Dimensions ............................................................................
Aft Cabin and Baggage Compartment .................................................
Fuselage Assembly ...............................................................................
Instrument Panel and Pedestal.............................................................
Overhead Console .................................................................................
Maintenance Ports .................................................................................
Caution and Warning Panel ..................................................................
Power Plant ............................................................................................
FADEC Control System Schematic ......................................................
Figure
Number
Page
Number
1-1 ...........
1-2 ...........
1-3 ...........
1-4 ...........
1-5 ...........
1-6 ...........
1-7 ...........
1-8 ...........
1-9 ...........
1-8
1-9
1-12
1-16
1-18
1-19
1-24
1-26
1-28
31 JAN 2007—Rev. 2———1-5
BHT-407-MD-1
MANUFACTURER’S DATA
FIGURES (CONT)
Subject
Auto to Manual Transition at Low Fuel Flow ......................................
Auto to Manual Transition at Intermediate Fuel Flow ........................
Auto to Manual Transition at High Fuel Flow......................................
Power Plant Components .....................................................................
Engine Oil System .................................................................................
Fuel System............................................................................................
Fuel Transfer System Schematic .........................................................
Left Fuel Boost/XFR Alternate Circuit Schematic ..............................
Transmission Assembly .......................................................................
Transmission Oil System......................................................................
Main Rotor Assembly ............................................................................
Tail Rotor Assembly ..............................................................................
Flight Controls .......................................................................................
Airspeed Actuated Pedal Stop System................................................
Hydraulic System...................................................................................
DC Electrical System.............................................................................
Pitot Static System ................................................................................
Figure
Number
Page
Number
1-10.........
1-11.........
1-12.........
1-13.........
1-14.........
1-15.........
1-16.........
1-17.........
1-18.........
1-19.........
1-20.........
1-21.........
1-22.........
1-23.........
1-24.........
1-25.........
1-26.........
1-36
1-38
1-40
1-58
1-62
1-70
1-71
1-75
1-80
1-82
1-84
1-85
1-88
1-91
1-92
1-93
1-98
Table
Number
Page
Number
1-1...........
1-2...........
1-42
1-49
1-3...........
1-4...........
1-5...........
1-52
1-65
1-66
1-6...........
1-67
1-7...........
1-67
1-8...........
1-67
1-9...........
1-10.........
1-67
1-78
TABLES
Subject
Time to Power Change ..........................................................................
250-C47B FADEC Software Version 5.202 Fault Code Display .........
250-C47B FADEC Software Version 5.356 (Reversionary Governor)
Fault Code Display ................................................................................
Engine Torque Exceedance Monitoring ..............................................
GAS PRODUCER GAUGE (NG) Exceedance Monitoring....................
MGT Exceedance Monitoring — During Start
(P/N 407-375-001-101/103).....................................................................
MGT Exceedance Monitoring — During Normal Operation
(P/N 407-375-001-101/103).....................................................................
MGT Exceedance Monitoring — During Start
(P/N 407-375-001-105 and Subsequent)...............................................
MGT Exceedance Monitoring — During Normal Operation
(P/N 407-375-001-105 and Subsequent)...............................................
Transfer Light Activation Table, S/N 53175, and Subsequent...........
1-6———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
Section 1
SYSTEM DESCRIPTION
1
1-1. INTRODUCTION
Be ll M od el 40 7 h elic op ter, p rima ry an d
auxiliary systems, and emergency equipment
are described within this section. Optional
equipment systems which do not require a
Flight Manual Supplement (FMS) will be
described in this section.
1-2. HELICOPTER DESCRIPTION
The Model 407 is a single engine, seven-place
light helicopter. Standard configuration
provides for one pilot and six passengers.
The fuselage consists of three main sections:
the Forward Section, the Intermediate
Section, and the Tailboom Section. The
forward section utilizes aluminum
honeycomb and carbon graphite structure
and provides the major load carrying
e l e m e n ts o f t h e f o r w a r d c a b i n . T h e
intermediate section is a semi-monocoque
structure which uses bulkheads, longerons
and carbon fibre composite side skins. The
ta i l b o o m i s a n a l u m i n i u m m o n o c o q u e
construction which transmits all stresses
through its external skins.
The helicopter is powered by a Rolls-Royce,
Model 250-C47B engine. Refer to paragraph
1-11, Power Plant, for complete system
description.
The main rotor is a four-bladed, soft-in-plane
design with a composite hub and individually
interchangeable blades. The tail rotor is a
two-bladed teetering rotor that provides
directional control. Refer to paragraph 1-28,
Rotor System, for a complete system
description.
Basic helicopter landing gear is the low skid
type. Optional pop-out emergency flotation
gear or high skid gear is also available.
1-3. PRINCIPAL DIMENSIONS
Principal exterior dimensions are shown in
Figure 1-1. All height dimensions must be
considered approximate due to variations in
loading and landing gear deflection. Principal
interior dimensions of the cabin and baggage
compartment are shown in Figure 1-2.
1-4. LOCATION REFERENCES
Locations on and within the helicopter can be
determined in relation to the fuselage
s tat io n s , w a te r lin e s , a n d b ut to c k lin e s
measured in inches (mm) from known
reference points.
1-4-A.
FUSELAGE STATIONS
Fuselage stations (FS or STA) are vertical
planes perpendicular to, and measured along,
the longitudinal axis of the helicopter. Station
zero is the reference datum plane and is 1.0
inch (25.4 mm) forward of the nose of the
helicopter or 55.16 inches (1401 mm) forward
of the forward jack point center line.
1-4-B.
WATERLINES
Wa t e r l i n e s ( W L ) a r e h o r i z o n ta l p l a n e s
perpendicular to, and measured along, the
vertical axis of the helicopter. Waterline zero
is a reference plane located 20.0 inches (508
mm) below the lowest point on the fuselage.
31 JAN 2007—Rev. 2———1-7
BHT-407-MD-1
NOTE
1
MANUFACTURER’S DATA
TYPICAL
When high gear is installed, dimension is 10.91 feet (3.33 m).
407_MD_01_0003
Figure 1-1. Principal Dimensions
1-8———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
20.75 IN.
(527.05 mm)
VIEW FROM LEFT-HAND SIDE
21.5 IN.
(546.1 mm)
42.0 IN.
(1067 mm)
37.0 IN.
(940 mm)
36.0 IN.
(914.4 mm)
AUX
FUEL
TANK
28.0 IN.
(711.2 mm)
16.5 IN.
(419 mm)
BAGGAGE COMPARTMENT FLOOR
VIEW LOOKING DOWN
WITH AUX TANK INSTALLED
TYPICAL
NOTES
1. Aft cabin space 85 cubic feet (2.4 m3)
Left forward cabin 20 cubic feet (0.6 m3)
Cabin floor stressed for 75 pounds per square foot (3.7 kg/100 cm2).
2. Baggage compartment 16 cubic feet (0.45 m3)
Baggage compartment floor stressed for 86 pounds per square foot (4.2 kg/100 cm2).
407_MD_01_0004
Figure 1-2. Aft Cabin and Baggage Compartment (Sheet 1 of 2)
31 JAN 2007—Rev. 2———1-9
BHT-407-MD-1
MANUFACTURER’S DATA
4
1
3
12
2
11
10
9
8
6
5
7
TYPICAL
VIEW FROM LEFT-HAND SIDE
1.
2.
3.
4.
5.
6.
39.3 inches (997 mm)
11.0 inches (279 mm) deep
36.5 inches (927 mm) forward side
12.0 inches (305 mm) high forward side
18.0 inches (457 mm)
46.0 inches (1168 mm)
7.
8.
9.
10.
11.
12.
49.0 inches (1245 mm)
15.5 inches (394 mm)
43.5 inches (1105 mm)
38.0 inches (965 mm)
58.5 inches (1486 mm) including litter door
33.5 inches (851 mm)
407_MD_01_0005
Figure 1-2. Aft Cabin and Baggage Compartment (Sheet 2 of 2)
1-10———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
1-4-C.
BUTTOCK LINES
Buttock lines (BL) are vertical planes
perpendicular to, and measured to the left and
right, along the lateral axis of the helicopter.
Buttock line zero is a plane at the lateral
centerline of the helicopter.
1-5. GENERAL ARRANGEMENT
The fuselage assembly (Figure 1-3) consists
of three main sections: the forward section
which extends from the cabin nose to the
bulkhead aft of the passenger compartment,
the intermediate section which extends from
t h e b u l k h e a d a ft o f t h e p a s s e n g e r
compartment to the tailboom attachment
bulkhead, and the tailboom section.
The forward section provides for pilot and
passenger seating, the fuel cell enclosures,
and the pylon support. The basic structure of
the forward fuselage utilizes two honeycomb
panels which are connected to create the
floor, and an aluminum roof beam assembly
which is attached to the center of another
honeycomb panel to create the roof. Two
bulkhead assemblies and a center post
connect the floor and roof to make an
integrated structure. The forward fuselage
section is closed out by two carbon fibre
composite side body fairings with
in t e r c h a n g e a b le c o m p o s i te d o o rs w it h
automotive type latches.
The landing gear is attached to the bottom of
the forward and aft bulkheads. The gear uses
a three point attachment configuration to
prevent ground resonance. The skid type
landing gear consists of two skids attached to
the ends of two arched crosstubes that are
secured to the fuselage by means of a three
point attachment configuration. Each skid
tube is fitted with a tow fitting, two saddles
with sockets for crosstubes, skid shoes along
BHT-407-MD-1
the bottom, a rear cap, and two eyebolt
fittings for mounting of ground handling gear.
The
intermediate
section
is
a
se m i-mo no co qu e s tru ctu re w h ich u s es
bulkheads, longerons, and carbon fibre
composite side skins which do not need
stiffeners or stringers. The lower section of
the intermediate fuselage is closed out by a
honeycomb composite fairing. The
intermediate section provides a deck for
engine installation, a baggage compartment,
and an equipment compartment under the
engine deck for electrical equipment, air
conditioning equipment, and the tail rotor
control servo actuator. The engine pan and
fo r w a rd a n d a ft f ir e w a lls o f t h e e n g in e
compartment are manufactured from titanium.
The tailboom is a monocoque construction
which transmits all stresses through its
external skins. To increase the strength of the
structure, certain skins incorporate bonded
doublers. The tailboom assembly includes the
tail rotor driveshaft, pitch change mechanism,
tail rotor gearbox, tail rotor hub and blade
assembly, and the horizontal stabilizer and
vertical fin.
Cowlings and fairings which enclose the
various roof and tailboom mounted
assemblies provide inspection access via the
use of hinge mounts, access doors, and
inspection windows or cutouts. They are
manufactured from either composite or
aluminum materials and are readily
removable for maintenance access.
The forward cowling is a composite
construction and incorporates a single point
hinge fitting which is located at the forward
end of the fairing. Two flush mounted toggle
hook latches secure the cowling. Support
rods are internally stowed on each side of the
co w l a nd f it in to ro of mo un ted c lips to
support the cowl in a raised position.
31 JAN 2007—Rev. 2———1-11
BHT-407-MD-1
MANUFACTURER’S DATA
10
9
8
12
6
5
11
13
7
14
4
19
3
2
20
21
22
1
23
25
24
15
18
16
17
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Windshield
Skylight window
Forward fairing
Transmission fairing assembly
Engine cowl
Aft fairing
Engine oil tank access door
Tail rotor driveshaft cover
Tail rotor gearbox fairing
Vertical fin
Tail skid and weight
Tailboom
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
Finlet
Horizontal stabilizer
Passenger door
Litter door
Crew door
Side body fairing
Oil cooler blower inlet duct
Aft skin panel
Air inlet cowl assembly
Baggage compartment
Forward fuselage
Lower window
Battery compartment door
407_MD_01_0006
Figure 1-3. Fuselage Assembly
1-12———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
The transmission cowling is a composite
construction and encloses the forward half of
the main transmission . This cow lin g is
mounted with fasteners. The fasteners are
contained in metal ejector clips which ensure
they are maintained with the cowling during
removal, and which ensure easy identification
of any fastener that may have dislodged from
its receptacle.
The air induction cowling is an aluminum
construction and encloses the aft half of the
main transmission. Inlet ducts are provided
on each side of the cowling to direct the
airflow into the induction screen or particle
separator. A small removable window is
provided on either side of the cowling to allow
checking of the inlet area when the particle
separator is installed. An inspection cutout is
provided on the right side to allow viewing of
the transmission oil level sight gauge. Two
hinged doors with flush-type latches are also
p r o v i d e d f o r i n s p e c t i o n o f t h e a ft
transmission area.
The engine cowling has composite hinged
side doors and an aluminum upper structure.
The side doors incorporate flush-type latches
and wing style stud fasteners. Attached to the
side doors are oil cooler blower inlet ducts
(S/N 53519 and subsequent or Post ASB
407-02-54). The doors are held in the open
position with mechanical support devices.
The side doors and upper structure of the
engine cowling incorporate screened vents to
allow air m ovement through the engine
compartment.
The aft cowling is a composite construction
and encloses the oil cooler and blower
assembly and the engine oil tank. It
incorporates a cutout to view the engine oil
tank sight gauge, and two doors which are
fastened with wing style stud fasteners. The
cowling incorporates screened cooling vents.
T h e ta i l b o o m u t i l i z e s t w o c o m p o s i t e
constructed fairings to enclose the tail rotor
driveshaft.
BHT-407-MD-1
The tail rotor gearbox fairings are both
composite constructed and secured with
screws. The upper and lower fairings both
incorporate hinge mounted doors for access.
1-5-A.
CREW COMPARTMENT
The crew compartment or cockpit occupies
the forward part of the cabin (Figure 1-3). The
pilot station is on the right side and the
copilot/forward passenger station is on the
left.
An instrument panel is mounted on a central
pedestal in front of crew seats. The panel is
tilted upward for maximum visibility from
either seat.
The overhead console is centered on the
forward cabin ceiling and incorporates most
of the electrical systems circuit breakers and
switches.
Each crew seat is covered with
flame-retardant fabric and is equipped with a
lap seat belt and a dual shoulder harness.
Each shoulder harness contains an inertia
reel which locks in the event of a rapid
deceleration.
A door on either side permits direct access to
the crew compartment. Each door is equipped
with interior and exterior handles which
operate a positive bolt latching mechanism.
T hi s s ty le o f d o or la tc h in g m ec h a ni s m
ensures smooth and positive operation when
opening and closing the doors. Each exterior
door handle incorporates a lock. The door
windows are made of gray tinted acrylic
plastic and incorporate a lower forward
sliding window for ventilation.
The main windshields, lower cabin nose chin
bubbles, and upper cab in ro of skylight
windows are also made of gray tinted acrylic
plastic.
31 JAN 2007—Rev. 2———1-13
BHT-407-MD-1
1-5-B.
PASSENGER COMPARTMENT
The aft area of the cabin (Figure 1-2) contains
a space of 85 cubic feet (2.4 m 3 ) for the
carrying of passengers or internal cargo. The
cabin can be configured with utility, standard,
or corporate interior kits. Each kit will include,
but is not limited to, vacuum formed
polycarbonate panels, armrests, door and
sidewall magazine pockets, assist steps, door
pulls, carpet, and all trim and attaching tracks
and hardware.
Basic configuration includes two aft facing
and three forward facing seats. All seats are
covered with flame-retardant fabric and are
equipped with lap seat belts and shoulder
harnesses. The shoulder harnesses lock in
the event of a rapid deceleration.
A cargo restraint kit is available for the
installation of tie-down provisions. This kit
provides forward bulkhead tie-down
provisions using four shackle/eyebolt
assemblies and floor mounted provisions
using four anchor plates. These provisions
allow cargo to be secured with a tie-down
assembly. It is the responsibility of the pilot to
ensure the cargo is adequately secured and
uniformly distributed.
A door on either side permits direct access to
the passenger compartment. In addition, the
left passenger door is hinged on the litter
door, so that the two may be opened together
to aid in the loading of the helicopter. The
LITTER DOOR caution light will illuminate if
the litter door is not properly secured. Each
door is equipped with interior and exterior
h a n d le s , w h ic h o p e ra t e a p o s it iv e b o lt
latching mechanism. This style of door
latching mechanism ensures smooth and
positive operation when opening and closing
the doors. Each exterior door handle
incorporates a lock. All cabin windows are
1-14———Rev. 4—30 APR 2008
MANUFACTURER’S DATA
made of tinted acrylic plastic and the
passenger doors can incorporate a lower
forward sliding window for ventilation.
1-5-C.
BAGGAGE COMPARTMENT
The baggage compartment (Figure 1-2) is
located aft of the passenger compartment and
has a capacity of 16 cubic feet (0.45 m3). The
compartment can carry up to 250 pounds
(113.4 kg) of baggage or other cargo which
can be secured using a tie-down assembly
and the tie-down fittings provided. It is the
responsibility of the pilot to ensure the cargo
is adequately secured and uniformly
distributed.
Access to the compartment is provided by an
exte rior d oor on the left side of the aft
fuselage. The door is hinged at the forward
end and opens the full width and height of the
co mpa rtment. It is s ecu red by tw o
push-button latches and a keyed lock. The
BAGGAGE DOOR caution light will illuminate
if the door is not properly secured.
A 19.2 US gallon (72.7 L) fuel tank may also be
installed in the baggage compartment as an
optional kit (BHT-407-FMS-6).
1-5-D.
TAILBOOM
The tailboom consists of the tailboom and the
components it supports: tail rotor and drive
system, vertical fin, and horizontal stabilizer
(Figure 1-3).
The tail rotor drive system consists of a
driveshaft mounted along the top of the
tailboom and a gearbox, which reduces the
tail rotor driveshaft input speed from 6317 to
2500 RPM. The gearbox also changes the
direction of drive 90° to accommodate the tail
rotor for directional control.
BHT-407-MD-1
Th e ve rt ic al fi n, c o m p os e d p ri m ar ily of
aluminum and honeycomb construction,
provides directional (yaw) stability and is
mounted on the aft end of the tailboom on the
right side. It contains a top fairing to mount the
anticollision light and on the lower edge, a
rubber bumper, and tail skid to protect the tail
rotor and fin in the event of a tail low landing.
The fin sweeps back, both above and below
the tailboom. The leading edge is canted
outboard 9° to reduce the required amount of
tail rotor thrust during forward flight at cruise
speed.
The horizontal stabilizer extends through the
tailboom, has leading edge slats, and small
auxiliary vertical fins or “dynamic dihedrals”.
The main body of the horizontal stabilizer is a
one-piece aluminum honeycomb structure. It
is an inverted airfoil which provides a
downward resultant lift on the tailboom to
maintain the cabin in a nearly level attitude
throughout all cruise airspeeds and to
aerodynamically streamline the fuselage to
reduce drag. The leading edge slats are
designed to improve pitch stability during
climbs.
The small auxiliary vertical fins are located on
each end of the horizontal stabilizer. The
leading edges of the fins are both offset 5°
outboard of the helicopter centerline. This
improves the roll stability of the helicopter in
forward flight.
The tailboom utilizes two composite
constructed fairings to enclose the tail rotor
driveshaft. The cowlings are mounted with
studs and receptacles. The studs are
contained in metal ejector clips, which ensure
the studs are maintained with the cowlings
during removal, and which ensure the easy
identification of any stud that may have
dislodged from its receptacle.
The tail roto r g earb ox fairings a re b oth
composite constructed and secured with
screws. A hinged door with wing style stud
fasteners is incorporated on the top fairing for
ease of servicing the gearbox. A hinged door
MANUFACTURER’S DATA
is incorporated on the bottom of the fairing for
ease of access to the tail rotor gearbox chip
detector. The tail rotor gearbox oil level may
be viewed through the aft screened cooling
vent of the fairing.
1-6. INSTRUMENT PANEL,
CONSOLE, AND PEDESTAL
1-6-A.
INSTRUMENT PANEL
Figure 1-4 shows the instrument panel with
optional equipment installed. The instrument
panel provides vibration dampening and is
hinge mounted on a central pedestal in front of
the crew seats.
The instrument panel is tilted at a 5° angle for
maximum visibility. The flight instruments are
located on the right side of the panel, and the
systems instruments are in two rows to the left
of the flight instruments. The caution and
warning panel is mounted just below the
glareshield across the top of the instrument
panel.
1-6-B.
OVERHEAD CONSOLE AND
PEDESTAL
The overhead console (Figure 1-5) is centered
on the cabin ceiling and incorporates most of
the electrical systems’ circuit breakers and
switches. The forward section of the panel has
integral lighting controlled by the instrument
lighting rheostat knob.
The pedestal (Figure 1-4) extends from the
instrument panel downward and aft to the
cabin seat structure. This forms a mounting
platform for optional equipment such as the
audio panel, radios, and gyrocompass control
panel etc. The lower left side of the pedestal
( F i g u r e 1 - 6 ) c o n t a i n s t h e FA D E C / E C U
maintenance
button,
FA D E C / E C U
maintenance port, ENGINE INSTRUMENT
maintenance port, 28 VDC auxiliary receptacle,
and the M/R RADS maintenance port and GPS
data loader port.
31 JAN 2007
Rev. 2
1-15
BHT-407-MD-1
MANUFACTURER’S DATA
1
FLOAT
ARM
BAGGAGE
DOOR
L/FUEL
BOOST
R/FUEL
BOOST
FADEC
FAULT
FADEC
FAIL
ENGINE
CHIP
GEN FAIL
BATTERY
RLY
AUTO
RELIGHT
LITTER
DOOR
L/FUEL
XFR
R/FUEL
XFR
RESTART
FAULT
FADEC
DEGRADED
XMSN
CHIP
XMSN OIL
PRESS
XMSN OIL
TEMP
START
HEATER
OVERTEMP
FUEL
FILTER
FUEL
VALVE
FUEL
LOW
FADEC
MANUAL
T/R
CHIP
CHECK
INSTR
HYD
SYSTEM
FLOAT
TEST
ENGINE
ANTI-ICE
FLOAT
TEST
BATTERY
HOT
ENGINE
OUT
ENGINE
OVSPD
PEDAL
STOP
CYCLIC
CENTERING
RPM
C/W
LT TEST
2
ON
A
U
T
O
E
L
T
RESET
6
7
0
8
5
9
4
3
10
2
1
0
11
% X 10
12
CONTROL
T
E
S
T
AIRSPEED
NR
120
10
8
3
9
2
10
1
T
50
33
W
N
5
NAV
3
W
24
21
10
~C
E
0
HDG
S
5
X 10
0
15
5
12
0
3
GPS
OBS
FADEC
MODE
12
PSI
5
X 10
0
2
E
10
10
2
FADEC SOFTWARE VERSION 5.356
WITH DIRECT REVERSION TO
MANUAL INSTALLED, REFER TO
FLIGHT MANUAL FOR OPERATION.
1
ENGAGED
2 MIN TURN
15
20
~C
4
PEDAL STOP
PTT
6
PSI
4
TEST
TEST
33
15
6
5
4
3
PER MIN
UP
6
15
10
8
2
FT
AVOID CONT OPS 68.4% TO 87.1% NP
30
0
11
3
12
HORN
1000
KING
RPM
% X 10
FWD TANK
OVSPD
60
70
30
FADEC
5
GS
24
11
12
40
80
4
0
3
DOWN
R
9
6
30
21
FUEL QTY
20
% RPM
90
21
30
5
2
FUEL CAPACITY
BASIC 869 LBS
WITH AUX 1005 LBS
(JET A AT 15° C)
4
5
1
HDG
W
7
8
S
15
GS
24
7
6
QTY LBS
X100
NAV
10
110
100
1
0
NP
12
10
E
~C X 100
0
3
2
299
7
6
2
0
5
1
3
N
X 10
FEE
00
9
4
1
0
4
100
T
R E C
S
8
0
6
15
AMPS
1
T
FEE
1010
L
PU
FOR
QUICK
80
3
12
60
8
7
C
H
K
KNOTS
100
8
M
U
T
E
FEE
100
9
L
5
16
2
6
H
O
R
N
E
20
4
40
100
3
I
N
S
T
R
120
DAVTRON
FUEL
PSI
20
150
N
SELECT
L
C
D
T
VOLTS
33
O.A.T
AUTO
TYPICAL
FUEL
VALVE
MAN
OFF
THIS HELICOPTER MUST BE OPERATED IN
COMPLIANCE WITH THE OPERATING LIMITATIONS
SPECIFIED IN THE APPROVED FLIGHT MANUAL
8
ON
7
6
1. Engine ANTI-ICE segment added
at S/N 53095 and subsequent.
Spare segment on S/N 53000
through 53094.
2. ELT remote switch
3. KI-525A horizontal situation
indicator
4. Airspeed Actuated Pedal Stop
press-to-test switch/annunciator
5. NAV/GPS switch
6. KI-208 course deviation indicator
7. KI-227 ADF indicator
8. Placard will reflect FADEC
software version installed.
407_MD_01_0007
Figure 1-4. Instrument Panel and Pedestal (Sheet 1 of 2)
1-16
Rev. 2
31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
9
10
11
12
13
14
15
16
9. KLN 89B GPS receiver
10. KMA 24H-71 audio panel
11. KX-155 or KX-165 VHF
communication/navigation transceiver
12. KY-196A VHF communication transceiver
13. KT-70 or KT-76A ATC transponder
14. KR-87 ADF receiver
15. KA-51B slave control
16. ICS mode selector
407_MD_01_0008
Figure 1-4. Instrument Panel and Pedestal (Sheet 2 of 2)
31 JAN 2007—Rev. 2———1-17
BHT-407-MD-1
MANUFACTURER’S DATA
HEADPHONE JACKS
FWD
TYPICAL
407_MD_01_0009
Figure 1-5. Overhead Console
1-18———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
INSTALLED WITH
KLN 89B GPS KIT
FADEC/ECU
GROUND
MAINTENANCE
CONNECTOR
FADEC/ECU
MAINTENANCE
BUTTON
LEFT SIDE LOWER CONSOLE
407MM_76_0012+
Figure 1-6. Maintenance Ports
31 JAN 2007—Rev. 2———1-19
BHT-407-MD-1
MANUFACTURER’S DATA
1-7. ENGINE INSTRUMENTS
The propulsion instruments are electronic
instruments that are individually powered
through circuit breakers located on the 28
VDC bus of the helicopter. Each side of dual
display instruments such as ENGINE OIL
PRESSURE and TEMPERATURE are powered
through individual circuit breakers.
tachometer will be displayed by no pointer
movement when depressing the LCD TEST
button. Any failure during the BIT found on
to r q u e , M GT, a n d N G i n d ic a to r s w il l b e
recorded in non-volatile memory (NVM) of the
instrument. Instruments that successfully
complete the BIT will have the first LCD of the
indicator lit or will display the appropriate
indication.
The dial of the instruments are made up of a
series of single bar LCDs which are arranged
in a tr en d a rc . Th e la s t L C D t o t u rn o n
indicates the reading of the instrument. The
first bar indicates power on. Some indicators
also include a digital display to provide more
accuracy of the information displayed.
On the torque MGT and N G indicators, a
letter E will appear in the digital display when
the instrument has recorded an exceedance.
Pressing the INSTR CHK button will display
the value of the last exceedance recorded.
Releasing the button will clear the E until the
next power up of the instrument.
1-7-A.
1-7-A-2.
INSTRUMENT OPERATION
All of the propulsion instruments have a
built-in-test (BIT) capability.
1-7-A-1.
POWER−ON BIT
The power-on BIT starts when the power is
applied to the instrument. The power-on BIT
does an integrity check of the electronic
components of the indicator.
NOTE
With the “Ahlers” NR/NP gauge
installed, the NR and NP needles do
not self test during power-on.
On all indicators except the dual tachometer,
during the power-on BIT, the trend arc display
shows the full scale for 6 to 8 seconds.
Torque and NG digits display 8188.8. MGT and
fuel digits display 81888. NR and NP needles
of the dual tachometer move to 107 and 100%
(upper redline limit) respectively.
If an LCD indicator fails the BIT, the trend arc
and the digits, if applicable, will show blank. If
faults are detected in the dual tachometer
during the BIT, pointer movement will not
occur. With the “Ahlers” dual tachometer
installed, no movement during power-on is
n o r m a l . F a u l ts i n t h e “ A h l e r s ” d u a l
1-20———Rev. 4—30 APR 2008
COMMANDED BIT
The commanded BIT starts when the LCD
TEST button is pushed. The commanded BIT
turns on all the LCDs so that the pilot can
verify that all the LCDs are working. On the
dual tachometer indicator, the individual
Rotor and Turbine pointers are driven to
indicate their respective upper redline limits.
1-7-A-3.
EXCEEDANCE MONITORING
TORQUE, MGT, and NG instruments have an
ability to record exceedances. Exceedances
have been predefined at the design stage and
preprogrammed into the microprocessor in
each indicator.
E x c e e d a n c e s a r e d e f i n e d a s l i m i ts o f
operation above which there may be some
maintenance action required.
NOTE
Exceedance monitoring is provided
as an aid to determining required
maintenance action. Only the person
in command of the helicopter at the
time the exceedance occurs can
verify that the exceedance recorded
reflects the actual occurrence.
The above instruments have built in NVM that
allows them to store up to 50 exceedance
MANUFACTURER’S DATA
events. For each exceedance, the memory
records the date, duration, and peak value
during the exceedance. If exceedance events
recorded are greater than 50, the earliest
dated exceedance that was in memory is
deleted and the latest exceedance is added in
its place.
To provide advance notice to the pilot that an
exceedance is about to be recorded, these
indicators also have other preprogrammed
advisory points at which they will flash the
Tre nd AR C dis play. T he digital re ad out
display will not flash. Specific values are
discussed in applicable systems descriptions
in this section of manufacturer’s data.
When an advisory is displayed by the
instrument, the instrument will also turn on
the CHECK INSTR light on the caution panel.
If the pilot makes control inputs to reduce the
instrument readings below the advisory
values, the advisory will no longer be
displayed and the CHECK INSTR light will be
turned off. No exceedance will be recorded.
When the indicator exceeds specifically
preprogrammed values, an exceedance will
be recorded. When an exceedance is about to
be recorded, the CHECK INSTR light will be
turned on by the instrument. In addition, when
an exceedance is recorded, the letter E will be
displayed at the left digit on the digital
display.
The pilot can acknowledge the exceedance
and cause the peak value to be displayed on
the analog and digital display by pushing the
INSTR CHK button (Figure 1-4) on the
instrument panel. The exceedance will display
for a maximum of 11 seconds, less if the pilot
pushes on the INSTR CHK button for a shorter
period of time. If there have been
exceedances recorded on different indicators,
e a c h i n d i c a t o r w i l l d i s p l a y i ts l a s t
exceedance.
Once the pilot has pushed the INSTR CHK
button and the exceedance(s) has been
displayed, the E will disappear from the digital
display. The E will not display until the power
BHT-407-MD-1
is removed from and applied to the indicator
(such as by pulling its circuit breaker or
turning helicopter power off and on).
The last exceedance (E) will continue to
display each time the indicator is powered up
until the exceedance(s) is removed from the
NVM of the indicator using a computer with
maintenance
download
s o ft w a r e
(BHT-407-MM).
If an exceedance has not been detected by the
TORQUE, MGT, or N G indicator, the CHECK
INSTR light and the E on the indicator digital
display will not illuminate when the INSTR
CHK button is pressed.
1-8. CHECK INSTRUMENT (CHECK
INSTR) LIGHT
The CHECK INSTR light circuit is designed to
alert the pilot that either the TORQUE, MGT, or
N G indicator is about to or has detected an
exceedance.
When the indicators exceed preset values, the
indicator's trend ARC display will begin to
flash and the CHECK INSTR light will turn on.
The digital readout display will not flash. If an
E is displayed by an indicator, the CHECK
INSTR light will remain on until the pilot
acknowledges the exceedance by pushing the
INSTR CHK button (Figure 1-4).
1-9. CLOCK
The clock is included in a multifunction
indicator mounted in the upper left area of the
instrument panel. The indicator also displays
Outside Air Temperature (OAT) in °C or °F and
volts (DC).
The clock is a digital display, quartz crystal
chronometer.
The clock has a high contrast liquid crystal
display and a two button control system
designed to prevent accidental time setting
while selecting various functions. The clock
functions are as follows:
Universal Time (UT) in 24-hour format.
30 APR 2008—Rev. 4———1-21
BHT-407-MD-1
MANUFACTURER’S DATA
Local Time (LT) in 12-hour format
(24-hour format may be selected).
Elapsed Time (ET) count up timer to
maximum of 99 hours, 59 minutes.
Elapsed Time (ET) count down timer from
maximum 59 hours, 59 minutes.
Elapsed Time (ET) alarm
display) at zero count down.
(flashing
Flight Time (FT) count up to a maximum
of 99 hours, 59 minutes.
Flight Time (FT) alarm (flashing display)
at zero count down.
A line on the display will be positioned over
the letters indicating the function selected
(UT/LT/FT or ET).
The clock display receives its power from the
28 VDC bus through the OAT/V INSTR circuit
breaker. When the clock is not powered, the
CONTROL and SELECT buttons are disabled.
The display lighting is powered by 5 VDC
instrument lighting system.
When all electrical power is turned off, the
clock display disappears, but the crystal
timing reference continues to operate from
the power of a 1.5 volt penlight dry cell
clipped to the back of the clock case. The dry
cell is not charged by the helicopter electrical
systems, and it should be replaced annually
to ensure uninterrupted operation of the
timing reference.
1-9-A.
CLOCK OPERATION
The SELECT button selects what is to be
displayed and the CONTROL button controls
w h a t is d is p l a y e d . T h e S E L E C T b u tt o n
sequentially selects UT, LT, FT, ET and back to
UT. The CONTROL button starts and resets
the elapsed time when momentarily pressed.
The Flight Time is recorded based on 28 VDC
power input received when the helicopter
hourmeter is running (paragraph 1-39-B).
1-22———Rev. 4—30 APR 2008
1-9-A-1.
UNIVERSAL TIME SETTING
Select UT using the SELECT button. Press the
SELECT
and
CONTROL
buttons
simultaneously to enter the set mode. The
tens of hours digit will start flashing. The
CONTROL button controls the flashing digit
and each push of the button increments the
digit. Once the flashing digit is set, the
SELECT button selects the next digit to be
set, from left to right across the display. After
the last digit is set, a final press of the
SELECT button will exit the clock from the set
mode. The function indicator will resume
normal flashing to indicate the clock is
running.
1-9-A-2.
LOCAL TIME SETTING
Select LT using the SELECT button. Press the
SELECT
and
CONTROL
buttons
simultaneously to enter the set mode. The
tens of hours digit will start flashing. The
setting operation is the same as UT, except
that the minutes are already synchronized
with UT and cannot be set in LT.
1-9-A-3.
ELAPSED TIME COUNT UP
Select ET using the SELECT button. Pressing
t h e C O N T R O L b u t t o n w i l l s ta r t t h e E T
counting up in minutes and seconds until
reaching 59 minutes and 59 seconds. The ET
will then start counting in hours and minutes
up to 99 hours and 59 minutes. Pressing the
CONTROL button again will reset the ET to
zero.
1-9-A-4.
ELAPSED TIME COUNT DOWN
Select ET using the SELECT button. Press the
SELECT
and
CONTROL
buttons
simultaneously to enter the set mode. The
count down time can now be set. Entering the
time is identical to UT time setting. When the
time is entered and the last digit is no longer
flashing, the clock is ready to start the count
down. Momentarily pressing the CONTROL
button starts the count down. When the count
down reaches zero, the display will flash
and the ET counter will begin counting up.
MANUFACTURER’S DATA
Pressing either the SELECT or CONTROL
button will deactivate the flashing alarm.
Pressing the SELECT button will select UT
display or pressing the CONTROL button will
reset the ET counter to zero.
1-9-A-5.
FLIGHT TIME RESET
Select FT using the SELECT button. Press and
hold down the CONTROL button for 3 seconds
or until the display shows 99:59. Flight time
will be zeroed upon release of the CONTROL
button.
1-9-A-6.
FLIGHT TIME ALARM SET
Select FT using the SELECT button. Press the
SELECT
and
CONTROL
buttons
simultaneously to enter the set mode. The
countdown time can now be set. Entering the
time is identical to UT time setting. When the
time is entered and the last digit is no longer
flashing, the clock is ready to start the count
down.
The Flight Time setting has not been affected
by setting the Flight Time Alarm. When the
Flight Time begins recording in conjunction
with the hourmeter (paragraph 1-39-B), the
Flight Time alarm will begin its countdown.
1-9-A-7.
FLIGHT TIME ALARM DISPLAY
NOTE
When the Flight Time is equal to the
alarm time, the display will flash. If FT
is not being displayed at the time the
alarm becomes active, the clock will
automatically select FT for display.
Press either the SELECT or CONTROL button
to turn off the alarm. Flight time will remain
unchanged and will continue counting.
1-9-A-8.
TEST MODE
Hold the SELECT button in for 3 seconds and
the display will read 88:88. This indicates all
digits are functioning properly.
BHT-407-MD-1
1-10. CAUTION AND WARNING
SYSTEM
The main function of the caution and warning
system is to detect specified conditions and
provide a visual or visual/audio indication.
Detection is accomplished through the use of
monitoring circuits that, when actuated, cause
the annunciator lights to illuminate. Visual
indications are provided through a caution
and warning panel while audio indications are
provided through three separate warning
horns: ENGINE OUT (pulsating), LOW ROTOR
(continuous), and FADEC FAIL (chime tone)
(Figure 1-7).
The caution and warning panel comprises 36
in di vid u al a n nu n ci ato r p o sit ion s w hic h
provide a visual indication of cautions (amber
lights), warnings (red lights), and advisory
(green or white lights) conditions. Each
annunciator contains three lamps. Individual
lamp operation is provided in the event one or
more of the lamps burns out.
Of the 36 annunciators, 32 are illuminated with
a ground input (ground seeking), three are
illuminated with a 28 VDC bus input (positive
s e e ki n g ), a n d o n e is i llu m in a te d w i th a
combination ground and 28 VDC bus input
(ground/positive seeking).
The FLOAT TEST, BATTERY HOT, ENGINE
OVSPD, ENGINE OUT, and RPM cannot be
d i m m e d . T h e r e m a i n i n g l i g h ts m a y b e
illuminated in either the fixed bright or the
fixed dim mode. With the instrument light
rheostat knob in the OFF position, all lights
will illuminate in the bright mode if they are
turned on. With the instrument light rheostat
knob, set between the dim and BRT position,
the CAUT LT DIM/BRT switch may be used to
select either the dim or bright mode. In the DIM
mode all lights, with the exception of those
mentioned above, will be illuminated at a fixed
dim value if they are turned on.
30 APR 2008
Rev. 4
1-23
BHT-407-MD-1
MANUFACTURER’S DATA
1
2
NOTES
1
Segment lettering for spare is
.
2
Engine anti-ice segment added at S/N 53095 and subsequent. Spare segment on S/N 53000 through 53094.
407_MD_01_0010
Figure 1-7. Caution and Warning Panel
1-24
Rev. 2
31 JAN 2007
MANUFACTURER’S DATA
All of the light lamps can be tested at once by
pressing the CAUTION LT TEST switch on the
right-hand side of the caution panel. This will
also test the FADEC Mode switch lamps and
NAV/GPS, Quiet Mode lamps (if installed). It
will not test the pedal stop switch lamps. The
pedal stop switch must be pressed (press to
Test-PTT) to check its lamps.
1-11. POWER PLANT
The Rolls-Royce 250-C47B engine (Figure 1-8)
is a turboshaft engine featuring a free power
turbine. The gas generator is composed of a
single-stage, single entry centrifugal flow
compressor directly coupled to a two-stage
gas generator turbine. The power turbine is a
two-stage free turbine which is gas coupled to
the gas gene ra to r tu rb in e. The integral
reduction gearbox has multiple accessory
pads and a splined output shaft which mates
with the freewheel unit. The engine has a
single combustion chamber with single fuel
injection and igniter. The output shaft center
line is located below the center line of the
engine and the single exhaust outlet is
directed upward. The engine incorporates a
Full Authority Digital Engine Control (FADEC)
system. At takeoff power (100% instrument
torque), the engine will provide 674
horsepower at sea level, 31°C, basic inlet and
at maximum continuous power (93.5%
instrument torque), the engine will provide
630 horsepower at sea level, 19.5°C, basic
inlet.
To further clarify the 250-C47B's engine
horsepower, the engine dataplate identifies
the engine's rated horsepower as 650. This
rated horsepower is what Rolls-Royce
guarantees the engine will provide at sea level
at a specific fuel flow and MGT. This ensures
all 250-C47B engines sold to Bell Helicopter
meet or exceed the rated horsepower
specification. Therefore, the rated
horsepower shown on the engine dataplate is
not the maximum horsepower that the engine
will deliver when installed in the 407
helicopter.
BHT-407-MD-1
1-12. ENGINE CONTROLS —
FADEC SYSTEM
This paragraph addresses the operational
design of the Rolls-Royce 250-C47B FADEC
system, its relationship with airframe and
ro t or s ys te m s, po s s ib le s ys t em fa u lts,
troubleshooting, and training procedures. The
information provided reflects operation with
FADEC software version 5.202 and 5.356 with
direct reversion to manual system installed.
Although Flight Manual Emergency
Procedures are explained in this section, refer
to the BHT-407-FM-1, Section 3, for actual
flight operations.
1-12-A. FADEC SYSTEM
The FADEC system is designed to enhance
flight safety and reduce pilot workload as well
as p rov ide o ther impo rta nt be ne fits . In
addition to the operational benefit of
increased TBO, engine automatic start, and
precise control of main rotor speed, the 407
features redundant signal sensing,
continuous monitoring, and self diagnostics.
Sinc e muc h o f the re dun dant desig n is
transparent to the pilot, the caution/warning/
advisory system has been expanded to advise
of conditions resulting from the increased
monitoring.
Although the possibility of a FADEC system
failure is unlikely, pilots and maintenance
personnel must have an operational
understanding of the FADEC system, along
with a sound knowledge of emergency and
troubleshooting procedures. Bell Helicopter
recommends that personnel involved with the
407 familiarize themselves with the procedure
for FADEC FAILURE. Familiarization with this
procedure will help to emphasize Flight
Manual Emergency Procedures as the
following FADEC information is read within
Manufacturer’s Data.
A FADEC FAILURE condition during operation
in AUTO mode will be indicated by activation
o f t h e FA D E C FA I L w a r n i n g h o r n a n d
illumination of the FADEC FAIL and FADEC
MANUAL warning lights.
30 APR 2008—Rev. 4———1-25
BHT-407-MD-1
MANUFACTURER’S DATA
3
2
4
1
5
TYPICAL
1.
2.
3.
4.
5.
Air inlet
Compressor assembly
Combustion section
Turbine section
Power and accessories gearbox
407_MD_01_0002
Figure 1-8. Power Plant
1-26———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
Respond to the horn and lights as described
in Section 3, Emergency/Malfunction
Procedures, in basic Flight Manual.
1-12-B. FADEC SYSTEM — OPERATION
The control system (Figure 1-9) is designed to
operate in many modes, covering all possible
system states, including the transitional
modes to and from the AUTO and MANUAL
modes.
The FADEC uses a single channel control with
one microprocessor and one electronic lane.
There is also a MANUAL mode hydro
mechanical back-up. The FADEC system has
two main components: the airframe mounted
Electronic Control Unit (ECU) and the engine
mounted Hydro mechanical Unit (HMU).
The ECU monitors numerous internal and
external inputs to modulate fuel flow and
therefore control engine speed, acceleration
rate, temperature, and other engine
parameters. The ECU provides inputs to the
HMU to modulate fuel flow based on the
continuous monitoring of the following:
Me asure d G as Tempe rature (M GT), G as
Producer speed (N G ), Power Turbine speed
(NP ), Main Rotor speed (NR), Engine Torque
Meter Oil Pressure (TMOP), Collective Pitch
(CP) and rate, Compressor Inlet Temperature
(CIT), Ambient Pressure (P1), and Power
Lever Angle (PLA)/throttle position.
In addition, when 5.356 software is installed,
the engine uses a digital electronic control
system based on two electronic governors
called primary channel and reversionary
g o v e r n o r. T h e p r i m a r y c h a n n e l i s a
fu ll-a uth or ity d ig ita l ele c tro nic co nt rol
(FADEC) that controls, monitors, and limits
engine power while maintaining helicopter
rotor speed. The reversionary governor can
automatically take control of the engine in the
event of a primary c hannel failure. The
reversionary governor uses a limited set of
i n p u ts a n d p r o v i d e s b a s i c e l e c t r o n i c
governing.
BHT-407-MD-1
The HMU consists of a two stage suction fuel
pump, fuel metering assembly, Auto/Manual
changeover solenoid valve, electric
overspeed valve, mechanical fuel shut off
valve, hot start fuel solenoid valve, and
altitude compensated bellows. The HMU
provides fuel modulation via a stepper motor
in A UTO mod e and a h ydro mec hanical
actuator in MANUAL mode.
1-12-C. POWER UP MODE AND
BUILT-IN-TEST
The FADEC system incorporates logic and
circuitry to perform self-diagnostics. In
general, sensors are checked for continuity,
rate and proper range. Discrete inputs are
checked for continuity and output drivers are
monitored for current demand to sense failed
actuators and open or shorted circuits. A
FADEC power up check exercises output
drivers and actuators to ensure system
functionality and readiness. The brief
appearances of light indications and their
respective horn observed immediately after
application of power are normal and are part
of the FADEC system’s designed initialization
process. If any faults are detected during the
self-test, the appropriate FADEC caution
panel light will illuminate.
The helicopter 28 VDC bus supplies electrical
power to the FADEC ECU until the engine
achieves 85% N P. Above this speed, the
FADEC ECU will select between the 28 VDC
bus and the engine-driven permanent magnet
alternator (PMA), as its primary power source.
The higher voltage source will be selected. In
the event of a primary power source failure,
the alternate source will be selected.
1-12-D. START IN AUTO MODE
To ready the system for an automatic start,
the FADEC MODE switch must be set to
AUTO, and the throttle set to the idle position.
T h e s ta r t s w i t c h i s t h e n m o m e n t a r i l y
positioned to START. Observe START and
AUTORELIGHT lights are illuminated before
releasing START switch. Throttle modulation
of fuel flow is not required.
26 MAR 2007—Rev. 3———1-27
BHT-407-MD-1
MANUFACTURER’S DATA
COCKPIT TYPICAL
FUEL NOZZLE
AIRFRAME
60
FUEL IN
POWER
LEVER
THROTTLE
(PLA) LINKAGE
FUEL
OUT
ELECTRONIC
CONTROL
UNIT (ECU)
VIBRATION
ISOLATORS
NG
60
120
HYDROMECHANICAL
UNIT
(HMU)
120
ENGINE
AMBIENT
PRESSURE
(P1)
NP /N R
MGT
ON
NG
OVERSPEED
TEST SWITCH
IGNITER RELAY
AUTO RELIGHT
FUEL
FILTER
NP
AUTO
(PMA)
PERMANENT MAGNET
ALTERNATOR
(CIT)
COMPRESSOR
INLET
TEMPERATURE
MAIN ROTOR
SPEED (N R)
(MGT)
MEASURED
GAS
TEMPERATURE
FADEC/ECU MTCE
PORT (EMC-35A
MTCE TERMINAL
CONNECTOR)
+28 VDC BATTERY/
AIRFRAME POWER
CAUTION
PANEL LIGHTS
-FADEC FAIL
-FADEC MANUAL
-FADEC DEGRADED
-FADEC FAULT
-RESTART FAULT
-ENGINE OVSPD
-ENGINE OUT
-AUTO RELIGHT
(MECHANICAL LINKAGE)
GAS GENERATOR
SPEED (N G)
POWER TURBINE
SPEED (N P )
MANUAL
FADEC MODE SWITCH
STARTER RELAY
LATCHING
THROTTLE
POWER LEVER
ANGLE (PLA)
TORQUE METER
OIL PRESSURE
(TMOP)
COLLECTIVE
PITCH (CP)
AND RATE
407MM_76_0001_c01+
Figure 1-9. FADEC Control System Schematic
1-28———Rev. 3—26 MAR 2007
MANUFACTURER’S DATA
Although the start sequence is automatic, the
pilot is responsible for monitoring the start
and taking appropriate action if required.
Therefore, it is recommended that both the
throttle and start switch are guarded until the
start is completed. Do not initiate a start if
FAD EC related cau tio n pane l lig hts are
illuminated unless appropriate maintenance
investigation or successful corrective action
has been carried out and no "current" faults
are shown (paragraph 1-12-R-2).
NOTE
After the throttle is set to idle, the
momentary contact start switch must
be activated within 60 seconds to
initiate the start and engage the
latching feature. The latching feature
of the start will engage when the
FADEC ECU senses momentary
activation (1 second) of the start
switch or upon sensing an NG speed
of 5%. If a start is attempted following
a delay of more than 60 seconds, the
FADEC system will not allow the
starter to latch following release of
the start switch, and will not
introduce fuel if the start switch is
held to START. Therefore, if a delay
of more than 60 seconds has
occurred, the system must be reset.
To reset the system, the throttle must
be repositioned to cutoff and then
back to idle. In addition, if electrical
power is interrupted prior to
initiating the start, with the throttle at
idle, the throttle must be
repositioned to cutoff and then back
to idle after power is restored to
re-enable the latching feature. A
normal automatic start sequence
may then commence. If starting
engine on external power, refer to
paragraph 1-32-A, External Power.
To allow for cooler starts and reduce the
possibility of reaching hot start abort limits, it
BHT-407-MD-1
is recommended that residual MGT be below
150°C, when below 10,000 feet HP or below
65°C, when above 10,000 feet HP prior to start.
To reduce residual MGT, a Dry Motoring Run
may be performed in accordance with the
Flight Manual.
Activating the automatic start mode engages
the airframe mounted FADEC/start relay
which is then latched by the FADEC ECU until
the NG speed reaches 50%. The FADEC/start
relay places the MGT indicator into the start
mode, signals the generator control unit/
voltage regulator to inhibit generator output,
flashes the shunt field for the duration of the
start, and activates the starter relay. The
starter relay activates the starter, illuminates
the START advisory segment on the caution
panel, and activates the igniter relay. The
igniter relay activates the engine igniter
system and illuminates the AUTO RELIGHT
advisory segment of the caution panel. While
i n s t a r t m o d e , t h e M GT i n d i c a t o r i s
p r o g r a m m e d t o r e c o r d s ta r t M GT
exceedances, should they occur, which differ
from normal operational MGT exceedance
limits.
NOTE
The following paragraph is
applicable if helicopter is configured
with FADEC Software Version 5.356
(Reversionary Governor).
Once N G speed reaches 10% for ambient
temperatures of 26.6°C (80°F) or below, or
12% for ambient temperatures above 26.6°C
(80°F), the FADEC system will introduce fuel,
detect the lightoff, and smoothly accelerate
the engine to id le while lim iting M GT if
necessary. At inlet temperatures below -18°C,
the FADEC will increase start acceleration
from 2 to 6% NG per second. The increase in
the acceleration rate will be noticeable during
start and improves start performance at
colder temperatures and at higher altitudes.
31 JAN 2007—Rev. 2———1-29
BHT-407-MD-1
MANUFACTURER’S DATA
NOTE
The following paragraph is
applicable if helicopter is configured
with FADEC Software Version 5.202.
Once N G speed reaches 10% for ambient
temperatures of -6.7°C (20°F) or below, or 12%
for ambient temperatures above -6.7°C (20°F),
the FADEC system will introduce fuel, detect
the lightoff, and smoothly accelerate the
engine to idle while limiting MGT if necessary.
At inlet temperatures below -18°C, the FADEC
will increase start acceleration from 2 to 6%
NG per second. The increase in the
acceleration rate will be noticeable during
start and improves start performance at
colder temperatures and at higher altitudes.
Upon reaching an engine N G speed of 50%,
the FADEC ECU unlatches the FADEC/start
relay, terminating the start sequence. In
a d d i t i o n , t h e E C U i n c r e m e n ts t h e
start-counter to the next number when lightoff
is detected.
Above 50% NG, the FADEC ECU carries out a
self test of the auto relight system and
continues to energize the igniter relay until 60
±1% NG. During this time, the AUTO RELIGHT
light will be illuminated. The engine will
con tin ue to a ccele rate u ntil reac hing a
stabilized idle of 63 ±1% NG.
The FADEC system also incorporates "Hot
Start Abort Logic" up to 50% NG during start.
This ECU controlled feature will cut off fuel
flow to the engine fuel nozzle if any of the
following conditions occur:
Start MGT exceeds 843°C, at pressure
altitudes less than 10,000 feet and if ECU
determined residual MGT was less than
82.2°C at initiation (5% NG) of start.
Start MGT exceeds 912°C, at pressure
altitudes greater than 10,000 feet or if
E C U d e t e r m i n e d r e s i d u a l M GT w a s
greater than 82.2°C at initiation (5% NG)
of start.
1-30———Rev. 2—31 JAN 2007
Voltage to FADEC ECU drops below 10.3
VDC. As a significant momentary voltage
drop occurs at initiation of the start,
ensuring a battery voltage of 24 VDC or
above prior to start, in conjunction with
appropriate battery maintenance, will
r e d u c e t h e p o s s i b i l i t y o f v o l ta g e
dropping to 10.3 VDC.
If a FADEC aborted start occurs or the pilot
manually initiates a start abort by positioning
the throttle to cut off, the FADEC is designed
to automatically keep the starter engaged for
up to 60 seconds from initiation of the start, to
reduce MGT to 150°C. Once 150°C MGT is
obtained, the starter will disengage.
NOTE
Momentarily positioning the start
switch to DISENG will only
deactivate the FADEC/start relay
which in turn disables the starter and
igniter circuits. In the event
deactivation of the starter and igniter
circuits occurs after engine light-off,
but below 50% NG, the FADEC ECU
will either modulate fuel flow to
provide a start if NG speed is
sufficient, or cutoff fuel flow if MGT
exceeds hot start abort limits due to
low NG speed. Therefore, positioning
the throttle to cutoff is the
appropriate method to manually stop
the start sequence.
Following start and prior to increasing engine
speed and main rotor RPM to 100%, a FADEC
MANUAL check is required. The purpose of
this check is to ensure the engine responds to
throttle movement while in MANUAL mode.
Prior to positioning the FADEC mode switch
to MANUAL for the purposes of this check,
ensure the throttle is positioned to idle and
NG speed is at 63 ±1%.
NOTE
The engine idle speed may reduce to
the point where the engine out light
and horn are activated.
MANUFACTURER’S DATA
NOTE
If NG increases to more than 75% NG
with throttle positioned in idle detent,
maintenance action is required.
Refer to the Rolls-Royce 250-C47B
Operation and Maintenance Manual.
Once the FADEC mode switch is positioned to
MANUAL, the FADEC MANUAL and AUTO
RELIGHT lights will illuminate. In addition, the
N G spe ed may change from 63 ±1%. An
increase or decrease in N G speed may be
noticed. The change in NG speed is due to the
f a c t t h a t t h e FA D E C s y s t e m i s n o t
temperature compensated in regard to fuel
flow in MANUAL mode. Engine load and HMU
calibration will also play a part in determining
if the NG speed will change.
Once NG is stabilized, slowly increase throttle
to approximately 5 to 10% N G to ensure
engine responds and then return throttle to
idle position. Position FADEC Mode switch to
AUTO and ensure FADEC MANUAL and AUTO
RELIGHT lights extinguish. Engine should
return to an NG speed of 63 ±1%.
Engine acceleration from idle to 100% NR/NP,
in AUTO mode, is achieved by smoothly
increasing the throttle to the FLY detent
position (PLA 70°). As the throttle is
positioned from idle (PLA 30° to 40°) to the
FLY detent position (PLA 70°), electrical
signals are sent to the ECU from the HMU –
PLA potentiometer. These signals dictate the
amount of authority the ECU has to control
maximum fuel flow (N G limiting) based on
throttle position, and in turn controls engine
N G s pe e d . T h e r e fo r e, a s t h e th r ot tl e is
increased from idle to the FLY detent position,
the fuel flow is electronically increased until
100% NR/NP is obtained. To avoid rapid engine
acceleration, it is recommended that throttle
ap plic atio n from idle to th e FLY d ete nt
position be conducted in a smooth and
gradual manner.
BHT-407-MD-1
1-12-E. ALTERNATE START — AUTO
MODE
For helicopters that operate at temperatures
of approximately 26.6°C (80°F) and above, or
at high altitudes and experience a hot start
abort event, the start procedure listed below
is approved and has been demonstrated to
overcome this issue in most instances.
For hot and/or high altitude environments,
and when prior troubleshooting has not
revealed any engine maintenance issues, this
pr oc e du re c an be us ed to a lle via te th e
possibility of aborted hot starts.
With the collective in the full down position
and the throttle in CUTOFF, the pilot can
initiate the start by pressing the starter
switch. This will energize the starter motor
and turn on the ignition exciter (continue to
hold starter switch). Once N 1 has reached
approximately 16%, the pilot is to move the
throttle from CUTOFF to IDLE. Once throttle is
moved to IDLE position, the starter will latch
and the starter switch can be released once
light off is detected. The engine will detect
light off and smoothly accelerate to ground
idle while limiting the MGT if necessary.
1-12-F. START IN MANUAL MODE
In accordance with the Rolls-Royce 250-C47B
Operation and Maintenance Manual, Manual
Mode starting on the ground is not authorized
except for use under emergency conditions or
under special permit from the local aviation
authority. Refer to the Rolls-Royce 250-C47B
Operation and Maintenance Manual to ensure
all Manual Mode Operational Procedures are
followed. Automatic hot start abort features
are not available in MANUAL mode.
To ready the system for a manual start, the
FADEC MODE switch is set to MAN and the
throttle positioned to cutoff. In MANUAL, the
FADEC will not latch the FADEC/start relay or
control the fuel scheduling during the start.
The start switch must be held in the START
position until the start sequence is completed
31 JAN 2007—Rev. 2———1-31
BHT-407-MD-1
MANUFACTURER’S DATA
at 50% NG. When the NG speed reaches 12%,
or NG speed specified in Rolls-Royce
250-C47B Operation and Maintenance Manual
for specific OAT, slowly advance the throttle
out of cutoff and stop when the engine lights
off. Allow the MGT to peak and then increase
fuel flow by modulating throttle to maintain
MGT within limits.
Once the engine has been started, position
the throttle to the idle detent and monitor the
NG speed.
The engine idle speed may reduce to
the point where the engine out light
and horn are activated.
NOTE
NOTE
If NG increases to more than 75% NG
with throttle positioned in idle detent,
maintenance action is required.
Refer to the Rolls-Royce 250-C47B
Operation and Maintenance Manual.
Idle detent speed in MANUAL may not
stabilize at 63 ±1% NG.
• If NG speed stabilizes below 63 ±1%,
adjust throttle to maintain 63 ±1% NG.
• If NG speed stabilizes between 63 ±1%
and 75% this is acceptable provided NP
speed does not fall into avoid STEADY
STATE range of 68.4 to 87.1% NP. If this
occurs, adjust throttle to avoid 68.4 to
87.1% NP.
NG,
O n c e a M A N U A L m o d e s ta r t h a s b e e n
initiated, it may be terminated at any time by
rotating the throttle to cutoff. The pilot should
continue motoring the engine until the MGT
h a s s ta b i l i z e d a t a n a c c e p ta b l e l e v e l .
Releasing the start switch from the START
position will disengage the FADEC/start relay
and disable the starter and igniter circuits.
1-32———Rev. 2—31 JAN 2007
Engine acceleration from idle to 100% is
achieved by positioning the throttle towards
full open. This is achieved through
hydromechanical control of the HMU fuel
metering valve.
1-12-G. IN-FLIGHT AUTO MODE
OPERATION
NOTE
• If NG stabilizes above 75%
maintenance action is required.
After a successful MANUAL mode start, and
idle has been achieved and is stable, perform
a momentary switch back to AUTO mode. If
engine flameout occurs, subsequent starts/
flight are prohibited until appropriate FADEC
system troubleshooting has been performed.
If throttle is not maintained in FLY
detent position during normal flight
operations, available engine power
may be limited. This will occur if
throttle is positioned from the FLY
detent (PLA 70°) to a setting which is
less than 62° PLA. Although any
throttle position from 62° PLA to full
open will provide the FADEC with
complete control to maintain N R
within limits, the approved throttle
position for flight is the FLY detent
position (PLA 70°).
During flight in AUTO mode with the throttle
in FLY detent position (PLA 70°), the FADEC
has complete control over engine operation to
maintain NR within limits. The ECU receives
engine and airframe parameter inputs and
cockpit command signals, processes them,
and modulates the HMU stepper motor driven
fuel metering valve to achieve desired engine
performance.
If required, as may be the case in certain
Emergency Procedures, an alternate means
of engine control is also available to the pilot
in AUTO mode. This can be achieved by
manipulating the throttle below FLY detent
position until the required engine
performance is achieved. As the throttle is
positioned between HMU Power Lever Angles
MANUFACTURER’S DATA
of 40 to 62°, electrical signals are sent to the
ECU from the HMU-PLA potentiometer. These
signals dictate the amount of authority the
ECU has to control maximum fuel flow (N G
limiting), and in turn, engine N G speed.
Therefore, as throttle is increased or
decreased, the maximum NG speed is
regulated electrically by limiting the fuel flow.
In AUTO mode, the FADEC is capable of
detecting an engine flameout by sensing NG
deceleration. Without any pilot action, the
auto-relight sequence is initiated by
es ta blis hin g a c on trol led fu el flow an d
activating the ignition system. The FADEC will
control the MGT and accelerate the engine to
the previously selected state.
The automatic relight sequence will initiate at
detection of a flameout, continuing the
procedure until a relight occurs or the N G
speed decays to 50%. During this time, the
AUTO RELIGHT and ENGINE OUT caution
panel lights will be illuminated and the engine
out warning horn will be activated. If the NG
decays below 50%, the FADEC system will
discontinue the attempt to relight the engine
and the ENGINE OUT light and warning horn
will remain active. In the event of an
unsuccessful relight, refer to the restart –
automatic mode emergency procedure
described in the BHT-407-FM-1. For additional
information on in-flight AUTO mode restarts,
refer to paragraph 1-18.
While in AUTO mode, if any failure occurs in
the ECU or in one of its input/output signals
that significantly impacts the ECU control of
the HMU, the pilot will be alerted via the
FADEC FAIL warning horn and the FADEC
FAIL and FADEC MANUAL warning lights. If
the detected failure does not significantly
impair the functioning of the ECU, the pilot
will be alerted via a FADEC DEGRADED,
FADEC FAULT caution light, RESTART FAULT
advisory light, or combination of, depending
on the nature of the fault.
BHT-407-MD-1
1-12-H. FADEC SYSTEM FAULTS
CAUTION
BELL HELICOPTER REQUIRES
MAINTENANCE ACTION, PRIOR TO
FLIGHT, WHEN A FADEC RELATED
LIGHT IS ILLUMINATED.
There are eight lights in the caution/warning/
advisory panel that are controlled by the
FA D E C : FA D E C FA IL , FA D E C M A N U A L ,
FADEC DEGRADED, FADEC FAULT, RESTART
FAULT, ENGINE OVSPD, ENGINE OUT, and
AUTO RELIGHT.
The FADEC ECU continuously monitors the
FA D E C s y s t e m f o r f a u l ts a n d m a k e s
appropriate accommodations to continue
operation. Fault codes have been
preassigned to those parameters being
monitored by the FADEC ECU.
Faults and exceedances can be recorded
under the following conditions:
Engine operating:
When the engine is operating (i.e.,
Iightoff has been detected) the
FADEC will automatically record
faults/exceedances as current, last
engine run, and accumulated, as
they occur.
Engine not operating:
When the engine is not operating, but
electrical power is applied, the
FADEC will only record current faults
as they occur.
If any failure occurs in the ECU/HMU or in one
of the input/output signals that significantly
impacts the ECU or control of the HMU, the
pilot will be alerted via the FADEC FAIL
warning horn and the FADEC FAIL/FADEC
MANUAL warning lights.
31 JAN 2007—Rev. 2———1-33
BHT-407-MD-1
With FADEC Software Version 5.356 installed,
the reversionary (backup) governor will be
activated under certain fault conditions to
eliminate a FADEC FAIL condition. This will
allow operations in a degraded mode while
remaining in AUTO mode (paragraph 1-12-M).
If the detected failure does not significantly
impair the functioning of the ECU, the pilot
will be alerted via a FADEC DEGRADED,
FADEC FAULT caution light, RESTART FAULT
advisory light, or a combination of, depending
upon the nature of the fault.
If the fault is minor in nature, it will not be
communicated to the pilot with the engine
ru n n i n g. T h e se f a u lts a re i d en t if ie d a s
maintenance advisory faults and will be
displayed during shutdown when the throttle
is placed in the cutoff position and NG speed
decays below 9.5%. This will be in the form of
a FADEC DEGRADED light.
MANUFACTURER’S DATA
w i t h a FA D E C FA I L / FA D E C M A N U A L
condition.
The DIRECT REVERSION to MANUAL system
ensures all FADEC failures revert directly to
MANUAL. FAIL FIXED failures do not exist
within this system. In addition, the system
incorporates a throttle which is detented at
the 90% NG bezel FLY position.
T he m ain in ten t o f D IR E C T to M A NU A L
system is to simplify pilot procedures in the
e v e n t o f a FA D E C f a i l u r e . T h i s i s
accomplished by allowing the pilot to keep
his hands on the controls during a FADEC
failure and enable an increase or decrease in
throttle from the FLY detent position, as
required. The pilot will only have to remove
his hand from the collective to press the
FADEC MODE switch to silence the horn once
established in MANUAL mode.
The BHT-407-FM-1 provides the appropriate
action required by the pilot for each light or
light/horn condition.
1-12-I-1.
All FADEC faults have been categorized into
five types. The first four relate to in-flight
faults and the fifth relates to Maintenance
Advisory faults with the engine shut down.
Maintenance Advisory faults displayed during
s h u td o w n w i l l b e d i s c u s s e d u n d e r t h e
heading FADEC SYSTEM FAULTS — ENGINE
SHUT DOWN.
In this situation, reversion to MANUAL mode
will occur, independent of the position of the
FADEC MODE switch on the instrument panel.
The reversion to MANUAL mode will begin
immediately.
1-12-I.
CATEGORY 1 — FADEC FAIL/
FA D E C M A N U A L — D I R E C T
REVERSION TO MANUAL
SYSTEM (DRTM)
F a u l ts t h a t r e q u i r e p i l o t a c t i o n a n d
automatically initiate a transition to the
MANUAL mode will be displayed immediately
when detected by the ECU. These faults will
display as FADEC FAIL/FADEC MANUAL. The
FADEC FAIL horn (chime tone) will activate in
conjunction with the FADEC FAIL and FADEC
MANUAL warning lights. In addition, the
RESTART FAULT light will also be displayed
1-34———Rev. 2—31 JAN 2007
THE FADEC SYSTEM WILL FAIL
TO THE MANUAL MODE AS
FOLLOWS:
This cond ition can be id entifie d by the
sounding of the FADEC FAIL warning horn,
illumination of the FADEC FAIL and FADEC
MANUAL caution panel lights and FADEC
MODE switch MAN light at the time of the
failure. As the FADEC SYSTEM has initiated
the transition to MANUAL mode, the pilot
must be aware that an increase or decrease in
N R /N P may occur within 2 to 7 seconds
following a direct failure to MANUAL. If this
occurs, collective and throttle will have to be
used to control RPM.
The objective of the DIRECT REVERSION to
MANUAL system is to provide the pilot with
throttle control of the fuel metering valve in a
timely manner. MANUAL mode allows the
MANUFACTURER’S DATA
pilot to control N R /N P with coordinated
control of the collective and throttle.
The following procedural steps are required:
1.
Throttle — If time permits, match
throttle bezel position to NG
indication.
This has proven to reduce the possibility
of an overspeed or underspeed
condition.
This procedure also permits a smoother
transition to MANUAL mode. This is due
to the fact that the actual NG speed and
fuel metering valve position prior to
switching to MANUAL mode will be very
close to that following the transition to
MANUAL mode, resulting in little, if any,
RPM change.
If time does not permit matching of
throttle bezel position to NG indication,
initially maintain throttle in detented FLY
position.
2.
NR/NP — Maintain 95 to 100% with
collective and throttle.
It is most important to ensure that N R /N P is
monitored and properly controlled, during
and following the transition to MANUAL
mode.
Within 2 to 7 seconds after FADEC FAIL
warning, NR/NP may increase or decrease
very rapidly. This will require collective
inputs to control RPM. The 407 rotor
system is very responsive to collective
inputs and can be controlled by the pilot
should a N R /N P overspeed/underspeed
tendency arise.
As it takes 2 to 7 seconds to complete the
transition to MANUAL mode, use of
BHT-407-MD-1
throttle to control NR/NP will be
ineffective until the transition to MANUAL
mode is complete. The transition will not
be completed until the fuel metering
valve in the HMU can be manually
controlled by the pilot through use of the
throttle on the collective.
There are two pistons within the HMU
(Figure 1-10): a Manual Load Piston (slow
piston) and a PLA Follower Piston (fast
piston), which must hydromechanically
extend to contact opposite sides of the
fuel metering valve shaft lever. The two
pistons move at different rates toward
the fuel metering valve lever. It takes
approximately 2.0 seconds for both
pistons to make contact with the fuel
metering valve lever following a
transition from an initial condition of low
fuel flow. Similarly, up to 7 seconds may
be required for the two pistons to make
contact following a transition from an
initial condition of high fuel flow. Refer to
Figure 1-10 through Figure 1-12 for
a d d i t i o n a l i n f o r m a ti o n o n A U TO t o
MANUAL mode transitions.
An increase in NR/NP speed may be
experienced while in transition to
MANUAL from a condition of low to
higher fuel flow or high fuel flow to a
higher fuel flow. This will be seen if the
throttle to bezel selection made by the
pilot in step 1 of the procedure is higher
than the actual N G speed at the time of
the FADEC FAILURE condition.
This will occur during the period when
the HM U M a nu al Lo a d Pis ton (slo w
piston) engages the fuel metering valve
l e v e r a n d m o v e s i t to a m o r e o p e n
position until the PLA Follower Piston
(fast piston) is contacted.
31 JAN 2007—Rev. 2———1-35
BHT-407-MD-1
MANUFACTURER’S DATA
1
AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
HIGH
PRESSURE
FUEL
VALVE
SHOWN
OPEN
(DE-ENERGIZED)
POWER
LOW
LEVER PRESSURE
FUEL
1.0 SECOND
2.0 SECONDS
MANUAL LOAD
PISTON (SLOW)
2 PLA FOLLOWER
PISTON (FAST)
METERING
VALVE
LEVER
MIN FUEL
FLOW
VARIABLE
ORIFICE
MAX FUEL
FLOW
PARTIAL CROSS SECTION OF HMU
407MM_76_0006+
Figure 1-10. Auto to Manual Transition at Low Fuel Flow (Sheet 1 of 2)
1-36———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
AUTO TO MANUAL TRANSITION AT LOW FUEL FLOW
INITIAL CONDITION:
Auto mode, low engine fuel flow, throttle at "FLY'' position.
After FADEC mode switch manual selection or initiation of direct reversion to manual:
Auto/manual changeover solenoid valve is de-energized, allowing high pressure fuel to manual
load piston and PLA follower piston.
Manual load piston "slowly'' extends. Engages metering valve lever in approximately 2.0 seconds
and begins to drive it to meet PLA follower piston.
Concurrently, PLA follower piston "rapidly'' extends in approximately 1.0 second to a position
that is a function of throttle (PLA) position.
Matching throttle and bezel to the actual N G speed will allow the PLA follower piston to position
itself very close to the actual position of the metering valve at the time of the transition. This will
minimize the fuel flow change during the transition. Positioning the throttle to a bezel setting that
is higher than actual NG speed at the time of the transition will produce an increase in fuel flow
during the transition. The increase in fuel flow will be caused as the manual load piston engages
the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the
throttle to a bezel setting that is lower than actual NG speed at the time of the transition will
produce a decrease in fuel flow during the transition. The decrease in fuel flow will be caused as
the PLA follower piston engages the metering valve lever and drives it towards the manual load
piston.
After both pistons engage, manual mode is established and no delay exists between throttle
(PLA) movement and fuel flow change.
Slew rate limiting is achieved by hydraulic dynamics.
NOTES
1
Auto/manual changeover solenoid valve normally closed (energized) in auto mode.
Auto/manual changeover solenoid valve is opened (de-energized) for transition to and
during manual mode operation. With the valve open, fuel pressure is used to position the
manual load piston and PLA follower piston.
2
PLA follower piston is controlled by throttle (PLA) position during transition to manual
and when in manual mode. PLA follower piston position is regulated by fuel pressure
bleed through variable orifice. This provides the means to increase or decrease fuel flow
by altering the position of the fuel metering valve. Manual load piston ensures metering
valve lever is held against PLA follower piston.
When in automatic mode, both the PLA follower piston and manual load piston are
retracted from the metering valve lever. They are held in the retracted position by fuel
pressure when the auto/manual changeover solenoid valve is closed (energized).
407MM_76_0007+
Figure 1-10. Auto to Manual Transition at Low Fuel Flow (Sheet 2 of 2)
31 JAN 2007—Rev. 2———1-37
BHT-407-MD-1
MANUFACTURER’S DATA
1
AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
HIGH
PRESSURE
FUEL
VALVE
SHOWN
OPEN
(DE-ENERGIZED)
POWER
LOW
LEVER PRESSURE
FUEL
0.5 SECOND
3 SECONDS
MANUAL LOAD
PISTON (SLOW)
2
METERING
VALVE
LEVER
MIN FUEL
FLOW
PLA FOLLOWER
PISTON (FAST)
VARIABLE
ORIFICE
MAX FUEL
FLOW
PARTIAL CROSS SECTION OF HMU
407MM_76_0008+
Figure 1-11. Auto to Manual Transition at Intermediate Fuel Flow (Sheet 1 of 2)
1-38———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
AUTO TO MANUAL TRANSITION AT INTERMEDIATE (CRUISE) FUEL FLOW
INITIAL CONDITION:
Auto mode, intermediate engine fuel flow, throttle at "FLY'' position.
After FADEC mode switch manual selection or initiation of direct reversion to manual:
Auto/manual changeover solenoid valve is de-energized allowing high pressure fuel to manual
load piston and PLA follower piston.
Manual load piston "slowly'' extends. Engages metering valve lever in approximately 3.0 seconds
and begins to drive it to meet PLA follower piston.
Concurrently, PLA follower piston "rapidly'' extends in approximately 0.5 second to a position
that is a function of throttle (PLA) position.
Matching throttle and bezel to the actual N G speed will allow the PLA follower piston to position
itself very close to the actual position of the metering valve at the time of the transition. This will
minimize the fuel flow change during the transition. Positioning the throttle to a bezel setting that
is higher than actual NG speed at the time of the transition will produce an increase in fuel flow
during the transition. The increase in fuel flow will be caused as the manual load piston engages
the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the
throttle to a bezel setting that is lower than actual NG speed at the time of the transition will
produce a decrease in fuel flow during the transition. The decrease in fuel flow will be caused as
the PLA follower piston engages the metering valve lever and drives it towards the manual load
piston.
After both pistons engage, manual mode is established and no delay exists between throttle
(PLA) movement and fuel flow change.
Slew rate limiting is achieved by hydraulic dynamics.
NOTES
1
Auto/manual changeover solenoid valve normally closed (energized) in auto mode.
Auto/manual changeover solenoid valve is opened (de-energized) for transition to and
during manual mode operation. With the valve open, fuel pressure is used to position the
manual load piston and PLA follower piston.
2
PLA follower piston is controlled by throttle (PLA) position during transition to manual
and when in manual mode. PLA follower piston position is regulated by fuel pressure
bleed through variable orifice. This provides the means to increase or decrease fuel flow
by altering the position of the fuel metering valve. Manual load piston ensures metering
valve lever is held against PLA follower piston.
When in automatic mode, both the PLA follower piston and manual load piston are
retracted from the metering valve lever. They are held in the retracted position by fuel
pressure when the auto/manual changeover solenoid valve is closed (energized).
407MM_76_0009+
Figure 1-11. Auto to Manual Transition at Intermediate Fuel Flow (Sheet 2 of 2)
31 JAN 2007—Rev. 2———1-39
BHT-407-MD-1
MANUFACTURER’S DATA
1
AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
HIGH
PRESSURE
FUEL
VALVE
SHOWN
OPEN
(DE-ENERGIZED)
POWER
LOW
LEVER PRESSURE
FUEL
0.1 SECOND
6 TO 7 SECONDS
MANUAL LOAD
PISTON (SLOW)
2
METERING
VALVE
LEVER
MIN FUEL
FLOW
PLA FOLLOWER
PISTON (FAST)
VARIABLE
ORIFICE
MAX FUEL
FLOW
PARTIAL CROSS SECTION OF HMU
407MM_76_0010+
Figure 1-12. Auto to Manual Transition at High Fuel Flow (Sheet 1 of 2)
1-40———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
AUTO TO MANUAL TRANSITION AT HIGH FUEL FLOW
INITIAL CONDITION:
Auto mode, high engine fuel flow, throttle at "FLY'' position.
After FADEC mode switch manual selection or initiation of direct reversion to manual:
Auto/manual changeover solenoid valve is de-energized allowing high pressure fuel to manual load piston
and PLA follower piston.
Manual load piston "slowly'' extends. Engages metering valve lever in approximately 6 to 7 seconds and
begins to drive it to meet PLA follower piston.
Concurrently, PLA follower piston "rapidly'' extends in approximately 0.1 second to a position that is a
function of throttle (PLA) position.
Matching throttle and bezel to the actual NG speed will allow the PLA follower piston to position itself very
close to the actual position of the metering valve at the time of the transition. This will minimize the fuel
flow change during the transition. Positioning the throttle to a bezel setting that is higher than actual NG
speed at the time of the transition will produce an increase in fuel flow during the transition. The increase
in fuel flow will be caused as the manual load piston engages the metering valve lever and drives it
towards the PLA follower piston. Inversely, positioning the throttle to a bezel setting that is lower than
actual NG speed at the time of the transition will produce a decrease in fuel flow during the transition. The
decrease in fuel flow will be caused as the PLA follower piston engages the metering valve lever and
drives it towards the manual load piston.
After both pistons engage, manual mode is established and no delay exists between throttle (PLA)
movement and fuel flow change.
Slew rate limiting is achieved by hydraulic dynamics.
NOTES
1
Auto/manual changeover solenoid valve normally closed (energized) in auto mode.
Auto/manual changeover solenoid valve is opened (de-energized) for transition to and during manual
mode operation. With the valve open, fuel pressure is used to position the manual load piston and PLA
follower piston.
2
PLA follower piston is controlled by throttle (PLA) position during transition to manual and when in
manual mode. PLA follower piston position is regulated by fuel pressure bleed through variable orifice.
This provides the means to increase or decrease fuel flow by altering the position of the fuel metering
valve. Manual load piston ensures metering valve lever is held against PLA follower piston.
When in automatic mode, both the PLA follower piston and manual load piston are retracted from the
metering valve lever. They are held in the retracted position by fuel pressure when the auto/manual
changeover solenoid valve is closed (energized).
407MM_76_0011+
Figure 1-12. Auto to Manual Transition at High Fuel Flow (Sheet 2 of 2)
31 JAN 2007—Rev. 2———1-41
BHT-407-MD-1
MANUFACTURER’S DATA
Inversely, a decrease in NR/NP speed may
be experienced during the transition to
MANUAL from a condition of low to lower
fuel flow or high fuel flow to a lower fuel
flow. This will be seen if the throttle to
bezel selection made by the pilot in
step 1 of the procedure is lower than the
actual NG speed at the time of the FADEC
FAILU RE con dition . This w ill o ccur
during the period when the PLA Follower
Piston (fast piston) engages the fuel
metering valve lever and moves it to a
more closed position as dictated by
throttle to bezel position.
desired power as selected by throttle
position for the transition. As stated
previously, the degree of power change
can be minimized by matching the
throttle bezel to the actual indicated NG
speed in step 1 of the FADEC FAILURE
procedure. This permits the smoother
transition to MANUAL mode due to the
fact that the actual N G speed and fuel
metering valve position prior to
switching to MANUAL will be very close
to that following the transition to
MANUAL mode. This will result in little, if
any, power/RPM change.
The approximate time to detect a power
change during the transition to manual is
summarized in Table 1-1.
Once both pistons contact the lever on
t h e f u e l m e t e r i n g v a l v e s h a ft , t h e
transition to M ANU AL Mo de will be
complete. The pilot will have slew rate
limited control of the fuel metering valve
via throttle position without any delay.
In simpler terms, there will be a time
delay and possible change in engine
power while the system transitions to
MANUAL. The length of the delay and
degree of pow er change during the
transition depends on engine power at
the time of the FADEC failure and the
Throttle may now be used, in conjunction with
collective, to maintain rotor and engine RPM
within 95 to 100%.
Table 1-1: Time to Power Change
ENGINE POWER AT TIME OF
FADEC FAILURE
DESIRED POWER AS
SELECTED BY THROTTLE
POSITION
APPROX. TIME TO DETECT
POWER CHANGE DURING
TRANSITION TO MANUAL
LOW POWER
HIGHER POWER
2.0 SECONDS
LOW POWER
LOWER POWER
1.0 SECOND
HIGH POWER
HIGHER POWER
7.0 SECONDS
HIGH POWER
LOWER POWER
0.1 SECOND
Once in MANUAL mode, the pilot will have
complete control of N R /N P by flight control
manipulation and the throttle on the
collective. The fuel flow slew rate is
hydromechanically limited to provide proper
responsiveness for helicopter operation and
also prevents surge. Fuel flow will be a
function of the pilot controlled fuel metering
valve orifice size. Maximum Continuous
P o w e r w ill b e a va i la b le f o r a l l a m b i en t
1-42———Rev. 2—31 JAN 2007
c on d iti on s . M A N U A L m od e fu el flo w is
pressure altitude compensated to maintain an
approximate constant horsepower, with a
change in altitude without throttle adjustment
by the pilot. Fuel flow in the MANUAL mode,
however, is not temperature compensated.
Because of this, there may be temperatures at
which maximum fuel flow in MANUAL mode
will not be sufficient to achieve Takeoff
Power.
MANUFACTURER’S DATA
NOTE
In the event engine (NP) overspeed
system is activated during transition
to or operation in MANUAL mode, the
control system is designed to keep
the engine running. Engine may
oscillate between 112.5 and 118.5%
N P until corrective action is taken
with throttle and collective
(paragraph 1-12-N).
3.
FADEC MODE switch – Depress one
time.
Depressing the FADEC mode switch one
time will mute the FADEC FAIL warning
horn (chime tone). As the transition to
MAN UAL mode was initiated by the
FADEC system, this step should not be
accomplished until pilot is firmly
established in MANUAL control.
This will allow pilot to keep hands on
flight controls during the transition to
MANUAL mode.
If the FADEC ECU is operational, it will
track HMU operation, perform
diagnostics, monitor engine functions,
and provide overspeed limiting for both
N P a n d N G. S u r g e d e t e c t i o n a n d
avoidance will not be available. If an
engine surge is encountered, decrease
the throttle until the surge condition
clears, then slowly increase the throttle
to the desired power level. Rapid power
changes should be avoided.
4.
Land as soon as practical.
BHT-407-MD-1
1-12-J. CATEGORY 2 — FADEC
DEGRADED
FADEC DEGRADED faults represent a loss of
some feature of the FADEC system which may
cause a degradation in performance. This
may result in N R droop, N R lag, or reduced
maximum power capability. These faults will
be displayed immediately when detected by
the ECU. Operations should be continued in
AUTO mode and helicopter is to be flown
s m o o t h l y a n d n o n a g g r e s s i v e l y. I n
conjunction with the FADEC DEGRADED
light, the RESTART FAULT light may also
activate under certain fault conditions.
With FADEC Software Version 5.356 installed,
the reversionary (backup) governor will be
activated under certain fault conditions,
which will also allow operations in a degraded
mode (paragraph 1-12-M).
Applicable maintenance action will be
required prior to next flight.
1-12-K. CATEGORY 3 — FADEC FAULT
FADEC FAULT indicates that PMA and/or MGT,
NP, or NG automatic limiting circuit(s) may not
be functional. In conjunction with activation
of the FADEC FAULT light, the RESTART
FAULT light may also activate under certain
f a u l t c o n d i t i o n s . T h e s e f a u l ts w i l l b e
displayed immediately when detected by the
ECU. Operations should continue in AUTO
mode. If both lights (FADEC FAULT and
R E S TA R T FA U LT ) a r e i l l u m i n a t e d , t h i s
indicates the MGT automatic limiting circuit
(905°C in flight), may not be functional. The
pilot should follow the appropriate
procedures as set out in the BHT-407-FM-1.
Applicable maintenance action will be
required prior to next flight.
Applicable maintenance action will be
required prior to next flight.
5.
1-12-L. CATEGORY 4 — RESTART
FAULT
Normal shutdown if possible.
If normal shutdown can not be completed
by rolling throttle to closed position, fuel
shutoff valve can be positioned to off.
RESTART FAULT indicates a subsequent
automatic engine start may not be possible.
31 JAN 2007—Rev. 2———1-43
BHT-407-MD-1
The fault does not require immediate action
by the pilot and should not affect
p e r f o r m a n c e o f t h e h e l i c o p t e r. I t i s
recommended that the pilot plan the landing
s i t e a c c o r d i n g l y. T h e s e f a u l ts w i l l b e
displayed immediately when detected by the
ECU and displayed as RESTART FAULT. Do
n o t a t t e m p t a s u b s e q u e n t s ta r t u n t i l
applicable maintenance action has been
completed.
If the engine shutdown procedures are not
properly followed, the MANUAL mode pistons
may begin to engage during the shutdown.
The FADEC may then be unable to prevent a
hot start on the next start and will indicate a
RESTART FAULT to warn the pilot. An HMU
manual piston parking procedure will be
required, as described in paragraph 1-12-P
and the Rolls-Royce 250-C47B Operations
and Maintenance Manual.
1-12-M. FADEC REVERSIONARY
GOVERNOR (FADEC SOFTWARE
VERSION 5.356)
The FADEC Reversionary (backup) Governor
consists primarily of a backup channel which
is contained in the ECU and is isolated from
the primary governor by a firewall for EMI. It
provides basic power turbine speed
governing in the event of a hard fault
occurring in the primary governor. Failure of
the FADEC into Reversionary Governor mode
is indicated by the illumination of the
following three lights: FADEC FAULT, FADEC
DEGRADE, and RESTART FAULT. This type of
failure causes a degradation in performance
and can cause NR droop, NR lag, or reduced
maximu m p ow er capa bility. O pe ratio ns
should be continued in AUTO mode and
he lico pte r is to b e flow n s mo oth ly an d
non-aggressively. Applicable maintenance
action will be required prior to next flight
(paragraph 1-12-R).
1-44———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
1-12-N. ENGINE OVERSPEED
PROTECTION
NP overspeed limiting is available in both the
AUTO and MANUAL modes by independent
ana lo g c irc uits integra l to the EC U. N G
overspeed limiting is available in the AUTO
and MANUAL modes through software control
in the ECU. In the event of a FADEC FAILURE,
it is possible that N G overspeed protection
will not be available.
The FADEC ECU continuously monitors for
N G and N P, or N P versus torque (Q) (5.202
FADEC software) overspeed conditions in
both AUTO and MANUAL Mode.
Activation of the ENGINE OVSPD warning
light will occur in the event of an NP
overspeed, N G overspeed, or if N P versus
torque (5.202 FADEC software) is above the
continuous limit (102.1% NP at 100% torque to
108.6% NP at 0% torque). With 5.356 FADEC
software, the ENGINE OVSPD warning light
will illuminate when N P reaches or exceeds
102.1%, regardless of the torque value. The
light will also momentarily illuminate during
the overspeed system test when the
overspeed solenoid valve closes.
If the ENGINE OVSPD light is activated during
engine operation due to an exceedance, and
the value has been recorded by the ECU, the
pilot will be provided with a maintenance
advisory on shutdown in the form of a FADEC
DEGRADED light. The FADEC DEGRADED
light will illuminate when N G speed decays
below 9.5%. If the pilot fails to recognize
illumination of the FADEC DEGRADED light
on shutdown, it will be illuminated the next
time electrical power is applied following the
FADEC system self test. When the FADEC
DEGRADED light is illuminated as a
m a i n t e n a n c e a d v i s o r y, m a i n t e n a n c e
investigation is required prior to further flight.
Peak values of exceedances are located on
the Engine History Data page of the EMC-35A
Maintenance Terminal.
MANUFACTURER’S DATA
If limits are exceeded, refer to 250-C47B
Series Overspeed Limits in the Rolls-Royce
Operation and Maintenance Manual and the
BHT-407-MM-2, Chapter 5.
• NP OVERSPEED
When the engine reaches 118.5 ±1% NP,
overspeed limiting will occur. The analog
overspeed limiting feature will activate
the overspeed solenoid valve, which
reduces fuel to the engine to a minimum
flow condition (sub-idle value of 34 to 45
pph). The minimum fuel flow increases
the likelihood of the engine remaining
running and recovering from the
overspeed. Once the NP speed drops to
112.5 ±2%, the overspeed solenoid valve
will be deactivated and fuel flow will
return to its previously commanded
value.
In the event the overspeed cannot be
controlled after fuel flow is reintroduced,
the overspeed limiting feature will control
the overspeed between the activation trip
p o i n t o f 11 8 . 5 ± 1 % N P a n d t h e
deactivation point of 112.5 ±2% NP. If this
occurs, attempt to control engine and
rotor speed with throttle and collective.
Refer to the ENGINE OVERSPEED
procedure in the Flight Manual.
• NG OVERSPEED
In AUTO mode, a software implemented
overspeed system is provided. Should
the software detect an NG overspeed, the
protection feature will be activated. In
addition, if the ECU has not failed, N G
overspeed protection will be available in
MANUAL mode.
When the engine reaches 110 ±1% NG,
the ENGINE OVSPD warning light will
illuminate and overspeed limiting will
occur. The software controlled overspeed
limiting feature will activate the
overspeed solenoid valve which reduces
fuel to the engine to a minimum flow
BHT-407-MD-1
condition (sub-idle value of 34 to 45 pph).
The minimum fuel flow increases the
l ik e l i h o o d o f th e e n g i n e re m a i n i n g
running and recovering from the
overspeed. Once the NG speed drops to
107 ±1%, the overspeed solenoid valve
will be deactivated and fuel flow will
return to its previously commanded
value.
In the event the overspeed cannot be
controlled after fuel flow is reintroduced,
the overspeed limiting feature will control
the overspeed between the activation
point of 110 ±1% NG and the deactivation
point of 107 ±1% N G. If this occurs,
attempt to control engine and rotor
speed with throttle and collective. Refer
to the ENGINE OVERSPEED procedure in
the BHT-407-FM-1.
• OVERSPEED SYSTEM SHUTDOWN TEST
Functionality of the overspeed system is
checked during FADEC power up and
thereafter continuously by the ECU.
Operation of the overspeed solenoid is
checked periodically by the pilot through
the use of the OVERSPEED SHUTDOWN
test procedure.
The OVERSPEED SHUTDOWN test
procedure will shut down the engine only
if collective pitch is below 10%, throttle
position is at idle, N G is between 60 to
66% and NP is less than 75%. The
OVERSPEED test button must be
pressed and held for a minimum of 1.0
second but not more than 10.0 seconds.
Once the test button is released, the
OVERSPEED test is completed as
follows. The FADEC ECU signals the
overspeed solenoid valve to close and
the ENGINE OVSPD light to come on.
Once the FADEC ECU senses an N G
decrease greater than 0.5%, the
overspeed solenoid valve is opened, the
ENGINE OVSPD light goes off, and the
engine is shut down by FADEC ECU
activation of the hot start abort feature.
31 JAN 2007—Rev. 2———1-45
BHT-407-MD-1
If the overspeed test is unsuccessful, the
engine will continue to operate at idle
power, the FADEC FAULT caution light
will illuminate, and a normal shutdown
procedure must be carried out.
1-12-O. ENGINE SHUTDOWN
Pilot control of engine speed from 100% NR/
NP to idle in AUTO mode is controlled through
throttle movement. As the throttle is
positioned from the detented FLY position to
idle, electrical signals are sent to the ECU
from the HMU – PLA potentiometer. These
signals dictate the amount of authority the
ECU has to control maximum fuel flow (N G
limiting), and in turn, engine speed. Therefore,
as throttle is decreased, the maximum fuel
flow that can be delivered to the engine is
reduced by positioning the fuel metering
valve to control engine NG speed/power.
In the unlikely event that a system fault
occurs which does not allow a reduction in
engine speed by positioning the throttle to
idle, complete the 2 minute cool-down at
100% flat pitch. After the 2 minute cool-down,
the engine is to be shutdown by rolling the
throttle to the CLOSED position.
Pilot control of engine speed from 100% NR/
NP to NG idle in MANUAL mode is controlled
hydromechanically through throttle
movement. Idle speed in MANUAL mode may
not stabilize at 63 ±1% N G. If this occurs,
maintain idle speed at 63 ±1% NG with throttle.
Following the appropriate cool down period at
idle, the engine may be shut down in either
the AUTO or MANUAL mode by positioning
the throttle to cutoff. This will close the
mechanical fuel shutoff valve within the HMU.
Do not reposition the throttle out of cutoff
unless NG has decayed to zero. If the throttle
is positioned out of cutoff prior to the N G
speed decreasing through 9.5%, the FADEC
IN-FLIGHT restart logic will introduce fuel and
activate the igniter. This can cause a relight
and possible over temperature condition
(paragraph 1-18).
1-46———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
If relight occurs, the pilot must immediately
position the throttle to the closed position
and activate the starter.
Additionally, the pilot must also allow NG to
de cay to 0% in AUTO mod e p rior to
positioning the battery switch to OFF. If this
procedure is not followed, the MANUAL mode
pistons may move hydromechanically after
electrical power is removed. This may cause a
RESTART FAULT the next time power is
applied to the FADEC ECU and a HMU manual
piston parking procedure will be required per
paragraph 1-12-P or as described in the
Rolls-Royce 250-C47B Operation and
Maintenance Manual.
1-12-P. HMU MANUAL PISTON PARKING
PROCEDURE
Starting with the HMU MANUAL mode pistons
in the wrong position may result in a hot start
of the engine. When the pilot is not certain of
the position of the pistons, or has received a
Maintenance Advisory that the pistons are out
of position, the following procedure will
assure the pistons are in the correct position
(fully retracted) for engine starting.
6.
Position throttle to cutoff.
7.
Pull IGNITER circuit breaker.
8.
BATT — ON.
9.
Power up check — Complete.
10. FADEC Mode switch — MANUAL.
11. Motor the engine (with throttle in
cutoff) for 10 seconds.
12. Wait for NG to decay to 0%.
13. FADEC Mode switch — AUTO.
14. Motor the engine (with throttle in
cutoff) for an additional 10 seconds.
15. Wait for NG to decay to 0%.
16. BATT — OFF.
17. Push in IGNITER circuit breaker.
MANUFACTURER’S DATA
BHT-407-MD-1
Continue prestart checklist.
1-12-Q. CATEGORY 5 — MAINTENANCE
A D V I S O R Y, FA D E C S Y S T E M
FAULTS — ENGINE SHUTDOWN
Maintenance Advisory Faults are those
detected by the ECU that are considered
minor in nature and are not communicated to
the cockpit with the engine operating. The
FA D E C D E G R A D E D li g h t s e r v e s a s t h e
maintenance advisory light. This light will be
illuminated, upon engine shutdown, if any
fault or exceedance has been detected during
the last engine run or if a current fault exists.
This will indicate that maintenance action is
required prior to the next flight.
Maintenance advisory faults will display
during shutdown when the throttle is placed
in the CLOSED position and the N G speed
decays below 9.5%. If the pilot misses the
maintenance advisory on shutdown, it will
illuminate at the next application of electrical
power.
1-12-R. CHECKING
CODES
FADEC
FAULT
As stated previously, faults can be displayed
im m e d i a te l y v ia a FA D E C FA I L , FA D E C
MANUAL, FADEC DEGRADED, FADEC FAULT,
R E S TA R T FA U LT o r b y a c o m b i n a t i o n
of these lights. Maintenance Advisory faults
will be displayed on shutdown via the FADEC
DEGRADED light. In addition, faults are also
used to identify exceedances.
Regardless if the fault light(s) were displayed
in-flight or at shutdown, maintenance action
is required prior to further flight. Refer to the
BHT-407-MM-9, Chapter 76 and Rolls-Royce
250-C47B Operation and Maintenance Manual.
The preferred method of determining FADEC
faults or exceedances is with the Goodrich
E M C - 3 5 A M a i n t e n a n c e Te r m i n a l . T h e
EMC-35A Maintenance Terminal (Windows
version) is capable of providing information
on Current Faults, Last Engine Run Faults,
Accumulated Faults (Fault History screen),
and N P overspeed exceedance information
(Engine History screen). Refer to the
Goodrich Maintenance Terminal User Guide
for operating instructions.
If a maintenance terminal is not available,
identify faults or exceedances using the
Maintenance Mode feature of the FADEC
system. This feature allows operators to
determine faults through a sequence of
flashing light displays on the cockpit caution
panel (paragraph 1-12-R-1).
1-12-R-1.
ENGINE RUN FAULT CODES —
PROCEDURE FOR VIEWING
The cau tio n pane l fault disp la y m ay be
operated as follows:
NOTE
Displayed faults may be LAST
ENGINE RUN or CURRENT faults.
Following this procedure, refer to
paragraph 1-12-R-2 to determine if
faults are current.
1.
Engine must be shut down and the
FADEC MODE switch positioned to
MANUAL. Place the collective full
down (below 10%) and the throttle in
the cutoff position.
NOTE
If the throttle or collective is moved
during the above procedure or the
FADEC MODE switch is positioned to
AUTO, the FADEC ECU will exit the
fault code reporting mode.
2.
Depress and release the FADEC ECU
maintenance button on the left hand
side of the lower pedestal to enter the
fault code reporting mode.
3.
FADEC DEGRADED, FADEC FAULT,
and the RESTART FAULT lights will
simultaneously flash five times to
31 JAN 2007—Rev. 2———1-47
BHT-407-MD-1
4.
MANUFACTURER’S DATA
indicate that maintenance mode has
been entered by the ECU.
(Reversionary Governor) Fault Code
Displays.
Depress and release the FADEC ECU
maintenance button. If a fault is
present, it will be displayed by a
specified
number
of
FADEC
DEGRADED caution panel light
segment flashes.
10. To determine fault description and
maintenance message code(s) from
caution panel flashing display, refer
to Table 1-2 or Table 1-3.
5.
Depress and release the FADEC/ECU
maintenance button to flash the next
fault code.
6.
Steady illumination of the FADEC
DEGRADED caution panel light
segment indicates that no other
faults exist for this light.
7.
Continue to depress and release the
FADEC/ECU maintenance button to
step through the FADEC FAULT and
RESTART FAULT caution panel light
segments, as above. This will
determine if fault codes exist for
these segments.
8.
9.
If no fault code exists for the selected
caution panel segment, the caution
light will illuminate continuously
when the FADEC/ECU maintenance
button is released.
When interrogation is complete, the
next
push
of
the
FADEC
maintenance button will cause the
FADEC DEGRADED, FADEC FAULT,
and RESTART FAULT caution light
segments to flash simultaneously
five times and then extinguish. This
indicates that the FADEC ECU has
exited the maintenance mode.
NOTE
Refer to Table 1-2 for FADEC
Software Version 5.202 and Table 1-3
for FADEC Software Version 5.356
1-48———Rev. 2—31 JAN 2007
1-12-R-2.
FADEC
FAULT
CODES
—
PROCEDURE TO DETERMINE
LA ST EN GIN E R U N F AU LTS
FROM CURRENT FAULTS
The procedure to determine if the fault codes
displayed are LAST ENGINE RUN faults or
CURRENT faults may be accomplished by
performing the steps listed in paragraph
1-12-R-1 with the throttle in the IDLE position.
With the throttle positioned to IDLE, any
FADEC fault code which is displayed will be a
current fault.
In addition, if no FADEC related lights are
displayed on the caution, warning, advisory
panel with electrical power applied, the
FA DE C in A UTO mo de, an d the th rottle
positioned to idle, no current faults exist.
1-12-R-3.
FAULT CODE CHARTS — USE OF
If fault codes have been displayed via the
caution panel, refer to Table 1-2 or Table 1-3
for specific information. This information can
be used in conjunction with the
B H T- 4 0 7 - M M - 9 , C h a p t e r 7 6 a n d t h e
Rolls-Royce Operation and Maintenance
Manual 250-C47B. These publications contain
the data to determine the maintenance action
required prior to further flight.
For specific fault information displayed on the
EMC-35A Maintenance Terminal, refer to the
Rolls-Royce 250-C47B Operation and
Maintenance Manual to determine the
maintenance action required prior to further
flight.
MANUFACTURER’S DATA
BHT-407-MD-1
Table 1-2: 250-C47B FADEC Software Version 5.202 Fault Code Display
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
FADEC DEGRADED cockpit
lamp flashes 1 time.
ECU Failure has occurred.
AD12bitFlt, AD8bitFlt, ECUOTFlt,
GainFlt, HLRfFLt, OffsFlt,
PROMFlt, PW10Flt,
RAMFlt, V15Flt, V5Flt, WDTFlt,
CJCFlt, OrDiodeFlt, BacCompFlt,
EEPROMFlt, ForCompFlt, P1Flt,
SWIntFlt, TestCelFlt, UARTFlt,
UUIntFlt, WDTOutFlt, OSVFlt,
NpOSFlt, OR28Flt, WDTTimeOut
FADEC DEGRADED cockpit
lamp flashes 2 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC DEGRADED cockpit
lamp flashes 3 times.
NP-Q Exceedance
NpQExLmAdv
FADEC DEGRADED cockpit
lamp flashes 4 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC DEGRADED cockpit
lamp flashes 5 times.
MGT indication failure
MGTFlt
FADEC DEGRADED cockpit
lamp flashes 6 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC DEGRADED cockpit
lamp flashes 7 times.
Failure to control HMU –
Auto/Manual Solenoid
AMSolFlt
FADEC DEGRADED cockpit
lamp flashes 8 times.
CIT temperature indication
failure
T1AFlt, T1BFlt, or T1ABFlt
FADEC DEGRADED cockpit
lamp flashes 9 times.
Metering Valve not in start
position
OpenMvFlg
FADEC DEGRADED cockpit
lamp flashes 10 times.
Starter Relay Interface
StrFlt
FADEC DEGRADED cockpit
lamp flashes 11 times.
NR Sensor – Rotor decay
anticipation
NrFlt
FADEC DEGRADED cockpit
lamp flashes 12 times.
Incorrect Overspeed Test
Switch indication
OSTstSwFlt
31 JAN 2007—Rev. 2———1-49
BHT-407-MD-1
MANUFACTURER’S DATA
Table 1-2: 250-C47B FADEC Software Version 5.202 Fault Code Display (Cont)
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
FADEC FAULT cockpit lamp
flashes 1 time.
HMU – Failure to control fuel
flow
WfLimFlag
FADEC FAULT cockpit lamp
flashes 2 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC FAULT cockpit lamp
flashes 3 times.
NP-Q Run Limit advisory
NpQRnLmAdv
FADEC FAULT cockpit lamp
flashes 4 times.
HMU Metering Valve
Potentiometer
WfMvFlt or WfStFlt
FADEC FAULT cockpit lamp
flashes 5 times.
Collective Pitch
Potentiometer indication
failure
CPFlt
FADEC FAULT cockpit lamp
flashes 6 times.
Failure to control HMU
(Stepper motor)
StepCntFlt, SMFlt, or AMSolFlt
FADEC FAULT cockpit lamp
flashes 7 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC FAULT cockpit lamp
flashes 8 times.
Failure to control HMU
(Overspeed Solenoid)
OSFlt
FADEC FAULT cockpit lamp
flashes 9 times.
Engine Surge event
SgFlag
FADEC FAULT cockpit lamp
flashes 10 times.
NG speed indication failure
Ng1Flt, Ng2Flt, or Ng12Flt
FADEC FAULT cockpit lamp
flashes 11 times.
Airframe Power Supply
failure
AF28Flt
FADEC FAULT cockpit lamp
flashes 12 times.
Engine Overspeed
OSFlag
RESTART FAULT cockpit
lamp flashes 1 time.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 2 times.
TMOP Sensor – Torque
indication failure
QFlt
1-50———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
Table 1-2: 250-C47B FADEC Software Version 5.202 Fault Code Display (Cont)
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
RESTART FAULT cockpit
lamp flashes 3 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 4 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 5 times.
Failure in HMU (PLA
Potentiometer) position
reading
PLA1Flt, PLA2Flt, PLA12Flt or
PLARfFlt
RESTART FAULT cockpit
lamp flashes 6 times.
PMA power supply failure
Al28Flt
RESTART FAULT cockpit
lamp flashes 7 times.
Failure to control HMU (Hot
Start Abort Solenoid)
StSFlt or StSIFlt
RESTART FAULT cockpit
lamp flashes 8 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 9 times.
Failure to control Ignition
Relay Interface
IgnFlt or IgnIFlt
RESTART FAULT cockpit
lamp flashes 10 times.
NP speed indication failure
Np1Flt, Np2Flt, or Np12Flt
RESTART FAULT cockpit
lamp flashes 11 times.
Incorrect Auto/Manual switch
indication
AMSwFlt
RESTART FAULT cockpit
lamp flashes 12 times.
Quiet Mode switch fault
QMSwFlt
FADEC DEGRADE, FADEC
FAULT and RESTART
FAULT cockpit lamps on
steady.
All faults have been
displayed.
31 JAN 2007—Rev. 2———1-51
BHT-407-MD-1
MANUFACTURER’S DATA
Table 1-3: 250-C47B FADEC Software Version 5.356 (Reversionary Governor)
Fault Code Display
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
FADEC DEGRADED
cockpit lamp flashes 1
time.
Primary Governor Failed
AD12bitFlt, AD8bitFlt, ECUOTFlt,
GainFlt, HLRfFLt, OffsFlt,
PROMFlt, PW10Flt, RAMFlt,
V15Flt, V5Flt, WDTFlt, CJCFlt,
OrDiodeFlt, BacCompFlt,
EEPROMFlt, ForCompFlt, P1Flt,
SWIntFlt, TestCelFlt, UARTFlt,
UUIntFlt, OSVFlt, NpOSFlt,
OR28Flt, WDTTimeOut,
ARINCFlt, ARINCHWFlt,
SWCfgFlt, RGSDFlt,QRawFlt,
T1BRawFlt, ESWRGFlt,
ESW2RGFlt, ESW3RGFlt,
ESW4RGFlt, ESW5RGFlt,
SWConfigRGFlt, PwrRstFlt,
RGSelSwFlt or NDOTWRCdRGFlt
FADEC DEGRADED
cockpit lamp flashes 2
times.
Reversionary Governor
Failed
AD10bitFltRG, PROMFltRG,
RAMFltRG, RGOTFltRG,
SWConfigFltRG, V10FltRG,
V15nFltRG, V15pFltRG,
V5qFltRG, WDTTimeOutRG,
ARINCHdFltRG, ARINCFltRG,
ARINCHWFltRG, BacCompFltRG,
ForCompFltRG, Or28FltRG,
OrDiodeFltRG, PW10LoFltRG,
RGTempFltRG, SPITempFltRG,
UARTFltRG, WDTFltRG,
PGSDHdFltRG, PGSDFltRG,
WfCorrPGHdFltRG,
WfCorrPGFltRG,
EngRnCtPGFltRG,
EngRnTmPGFltRG,
ESWPGFltRG, NpIncPGFltRG,
Np2RawPGFltRG,
P1RawPGFltRG,
T1ARawPGFltRG, SwPwrFltRG
FADEC DEGRADED
cockpit lamp flashes 3
times.
NP Exceedance
NpLmTOut
1-52———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
Table 1-3: 250-C47B FADEC Software Version 5.356 (Reversionary Governor)
Fault Code Display (Cont)
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
FADEC DEGRADED
cockpit lamp flashes 4
times.
Reversionary Governor did
not govern when Primary
Governor failed.
ECUGovFltRG
FADEC DEGRADED
cockpit lamp flashes 5
times.
MGT indication failure
MGTFlt
FADEC DEGRADED
cockpit lamp flashes 6
times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC DEGRADED
cockpit lamp flashes 7
times.
Failure to control Auto/
Manual Solenoid
AMSolFlt, AMSolFltRG
FADEC DEGRADED
cockpit lamp flashes 8
times.
CIT temperature indication
failure
T1AFlt, T1ABFlt, T1BFltRG or
T1DFltRG
FADEC DEGRADED
cockpit lamp flashes 9
times.
Metering Valve is not in start
position.
OpenMvFlg
FADEC DEGRADED
cockpit lamp flashes 10
times.
Starter Relay Interface
StrFlt
FADEC DEGRADED
cockpit lamp flashes 11
times.
NR Sensor – Rotor decay
anticipation
NrFlt
FADEC DEGRADED
cockpit lamp flashes 12
times.
Incorrect Overspeed Test
switch indication
OSTstSwFlt
FADEC FAULT cockpit
lamp flashes 1 time.
HMU – Failure to control fuel
flow
WfLimFlag, WfLimFlagRG
FADEC FAULT cockpit
lamp flashes 2 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC FAULT cockpit
lamp flashes 3 times.
NP Run Limit
NpRLmTOut
31 JAN 2007—Rev. 2———1-53
BHT-407-MD-1
MANUFACTURER’S DATA
Table 1-3: 250-C47B FADEC Software Version 5.356 (Reversionary Governor)
Fault Code Display (Cont)
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
FADEC FAULT cockpit
lamp flashes 4 times.
Failure in HMU Metering
Valve Potentiometer reading
WfMvFlt, WfStFlt or WfStFltRG
FADEC FAULT cockpit
lamp flashes 5 times.
Collective Pitch
Potentiometer indication
failure
CPFlt
FADEC FAULT cockpit
lamp flashes 6 times.
Failure to control HMU
(Stepper motor)
StepCntFlt, SmFlt, AMSolFlt or
SmFltRG
FADEC FAULT cockpit
lamp flashes 7 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
FADEC FAULT cockpit
lamp flashes 8 times.
Failure to control HMU
(Overspeed Solenoid)
OSFlt
FADEC FAULT cockpit
lamp flashes 9 times.
Engine Surge event
SgFlag
FADEC FAULT cockpit
lamp flashes 10 times.
NG speed indication failure
Ng1Flt, Ng2Flt, Ng12Flt or
Ng1FltRG
FADEC FAULT cockpit
lamp flashes 11 times.
Airframe Power failure
AF28Flt or AF28FltRG
FADEC FAULT cockpit
lamp flashes 12 times.
Engine Overspeed
OSFlag or OSEventLmpRG
RESTART FAULT cockpit
lamp flashes 1 time.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 2 times.
TMOP Sensor – Torque
indication failure
QFltRG
RESTART FAULT cockpit
lamp flashes 3 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 4 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
1-54———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
Table 1-3: 250-C47B FADEC Software Version 5.356 (Reversionary Governor)
Fault Code Display (Cont)
STATUS MESSAGE
FAULT DESCRIPTION
MAINTENANCE MESSAGE CODE
RESTART FAULT cockpit
lamp flashes 5 times.
Failure in HMU (PLA
Potentiometer) reading
PLA1Flt, PLA2Flt, PLA12Flt or
PLARfFlt
RESTART FAULT cockpit
lamp flashes 6 times.
PMA power supply
Al28Flt or Al28FltRG
RESTART FAULT cockpit
lamp flashes 7 times.
Failure to control HMU (Hot
Start Abort Solenoid)
StSFlt, StSIFlt, StSFltRG or
StSIFltRG
RESTART FAULT cockpit
lamp flashes 8 times.
This Status Message is not
used. Verify that message is
indicated by lamp.
Not Used
RESTART FAULT cockpit
lamp flashes 9 times.
Failure to control Ignition
Relay Interface
IgnFlt or IgnIFlt
RESTART FAULT cockpit
lamp flashes 10 times.
NP speed indication failure
Np1Flt, Np2Flt, NpDFlt, Np12Flt,
NpDFltRG or Np1FltRG
RESTART FAULT cockpit
lamp flashes 11 times.
Incorrect Auto/Manual
switch indication
AMSwFlt or AMSwFltRG
RESTART FAULT cockpit
lamp flashes 12 times.
Quiet Mode switch fault
QMSwFlt
FADEC DEGRADE, FADEC
FAULT, and RESTART
FAULT cockpit lamps on
steady.
All faults have been
displayed.
1-12-S. CLEARING
CODES
FADEC
FAULT
Faults/exceedances are not to be erased
unless appropriate maintenance actions have
been carried out in accordance with the
BHT-407-MM and Rolls-Royce 250-C47B
Operations and Maintenance Manual. Do not
attempt to clear any fault or exceedance while
the engine is operating.
If maintenance actions have been conducted
due to a recorded exceedance or to correct a
current or last engine run fault, it must be
ensured that no FADEC system lights are
illuminated when the throttle is positioned to
IDLE for the next start attempt. If a FADEC
related light is illuminated with the throttle
positioned to IDLE, a current fault exists and
further maintenance action is required.
1-12-S-1.
CURRENT FAULTS
Current faults may be cleared by performing a
power reset (battery switch OFF/ON). If fault is
n o l o ng e r d e te c te d , a s s oc ia t ed FA D E C
DEGRADED light will be extinguished.
31 JAN 2007—Rev. 2———1-55
BHT-407-MD-1
1-12-S-2.
LAST ENGINE
EXCEEDANCES
MANUFACTURER’S DATA
RUN
FAULTS/
LAST ENGINE RUN faults or exceedances
may be cleared by performing a successful
engine start or with use of the EMC-35A
Maintenance Terminal.
To e r a s e L A S T E N G I N E R U N f a u l ts /
exceedances, refer to the EMC-35A
M a i n t e n a n c e Te r m i n a l U s e r s G u i d e f o r
operating instructions.
1-12-S-3.
ACCUMULATED FAULTS/
EXCEEDANCES
Accumulated faults or exceedances may only
be cleared with use of the EMC-35A
Maintenance Terminal.
1-12-U. COCKPIT PROCEDURAL
TRAINING ON GROUND (ENGINE
NOT RUNNING)
1.
Pull START and IGNTR circuit
breakers and ensure fuel valve
switch is positioned to OFF.
2.
Connect external power source.
3.
Allow instruments and FADEC to
complete self test with FADEC MODE
switch in AUTO.
4.
Position throttle to FLY detent and
collective to approximate cruise
flight setting.
5.
Simulate FADEC FAILURE by pulling
FADEC circuit breaker. This will
simulate a failure direct to MANUAL
mode. FADEC FAIL warning horn will
activate along with illumination of
the FADEC FAIL and FADEC
MANUAL caution panel lights.
FADEC MODE switch will illuminate
MAN.
6.
Carry
out
the
appropriate
BHT-407-FM-1 emergency response
procedure. (Depressing the FADEC
MODE switch will silence the horn.)
7.
Push in FADEC circuit breaker and
set FADEC MODE switch to AUTO.
8.
Repeat procedure
understood.
9.
Disconnect external power source,
position throttle to cutoff and push in
START and IGNTR circuit breakers.
N P exceedance values may only be cleared
from engine history data page with use of the
EMC-35A Maintenance Terminal.
1-12-T. FADEC TRAINING IN MANUAL
MODE
Prior to actual flight training in the MANUAL
mode, an operational understanding of the
FADEC system along with a sound knowledge
of emergency procedures is required. It is
recommended that training be first
accomplished in the helicopter in a cockpit
procedural environment (engine not running)
(paragraph 1-12-U). This can provide a visual
simulation of the horn and lights associated
with a FADEC failure direct to MANUAL mode
and familiarize the pilot with the required
cockpit actions. This should be followed by a
takeoff/hover/circuit and landing in MANUAL
mode which will allow the pilot to become
familiar with the required manipulation of the
throttle and controls (paragraph 1-12-V).
Once the pilot is comfortable with flight in
MANUAL mode, simulated FADEC failure
emergency procedures can be carried out in
flight (paragraph 1-12-W).
1-56———Rev. 2—31 JAN 2007
until
it
is
1-12-V. FLIGHT TRAINING IN MANUAL
MODE
In MANUAL mode, the following is applicable:
• Switching to MANUAL mode at IDLE
(63 ±1% NG) may result in change in
NG.
• Igniter operates continuously.
MANUFACTURER’S DATA
• AUTO RELIGHT, FADEC MANUAL
caution panel lights illuminate.
• FADEC MODE switch indicates MAN.
• FADEC ECU remains operational.
However,
surge
protection
and
avoidance logic is not available. In the
event of a FADEC FAILURE while
training in MANUAL mode, the FADEC
FAIL warning light will illuminate.
FADEC FAIL horn will not sound.
Remain in MANUAL mode and land as
soon as practical.
• Maximum
Continuous
Power
is
available for all ambient conditions.
Takeoff Power may not be available.
1.
Perform START procedure and
SYSTEMS CHECKS in AUTO mode.
2.
At idle (63 ±1% NG), depress FADEC
MODE switch to transition to
MANUAL mode.
NOTE
Transition back to AUTO mode can
be made at any time by depressing
FADEC MODE switch to AUTO. Upon
selecting AUTO mode, an engine
power transient may be experienced
as the FADEC ECU matches engine
power to rotor load. Ensure throttle is
positioned to FLY detent position
following selection of AUTO mode.
3.
Manipulate throttle on ground to
become familiar with MANUAL
control.
4.
Increase throttle to maintain 95 to
100% NR/NP. Manipulate throttle/
collective and lift into hover.
5.
When comfortable with manipulation
of throttle and flight controls in
BHT-407-MD-1
hover, transition into forward flight
and conduct circuits to touchdown.
1-12-W. SIMULATED FADEC
TRAINING (IN-FLIGHT)
FAILURE
When comfortable with flying circuits to touch
down in MANUAL mode, transitions between
AUTO and MANUAL may be conducted in
flight.
Simulation can be accomplished by activating
the FADEC FAIL horn by pushing the FADEC
FAIL horn test button. Although the applicable
FA D E C c a u t i o n pa n e l l i g h ts c a n n o t b e
activated, the pilot should respond to the
FA D E C FA I L h o r n a n d c a r r y o u t t h e
B H T- 4 0 7 - F M - 1 e m e r g e n c y r e s p o n s e
procedure. Once the throttle bezel position is
matched to the actual N G indication, pilot
shall position the FADEC MODE switch to
MANUAL.
Following simulated FADEC FAILURE training
in flight, ensure FADEC MODE switch is
positioned to AUTO prior to shutdown. This
will ensure the manual pistons in the HMU are
parked, and a subsequent start in AUTO can
be carried out without maintenance action.
1-13. COMBINED ENGINE FILTER
ASSEMBLY (CEFA)
The CEFA (Figure 1-13) provides both fuel and
scavenge lube filtration within a single filter
assembly which consists of a fuel filter bowl,
fuel bypass valve, fuel differential pressure
indicator, manifold assembly, a disposable
fuel filter element, a lube filter bowl, lube
bypass valve, lube differential pressure
indicator, and a disposable lube filter element.
The CEFA is located on the lower left hand
portion of the power and accessory gearbox
assembly.
31 JAN 2007—Rev. 2———1-57
BHT-407-MD-1
MANUFACTURER’S DATA
COMPRESSOR
TURBINE
COMBUSTION SECTION
8
7
6
9
5
4
3
2
10
1
POWER ACCESSORIES GEAR BOX
TYPICAL
1.
2.
3.
4.
5.
Ignition exciter box (Ref)
Torque meter oil pressure port (Ref)
Engine oil outlet port (Ref)
Engine oil inlet port (Ref)
Engine oil pressure port (Ref)
6.
7.
8.
9.
10.
Engine intake (Ref)
Shroud bleed air manifold connection (Ref)
Oil filter assembly
Bleed air valve (acceleration)
CEFA fuel filter impending bypass button
407_MD_01_0011
Figure 1-13. Power Plant Components (Sheet 1 of 2)
1-58———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
13
12
14
11
15
10
16
23
24
17
22
21
20
19
18
TYPICAL
11.
12.
13.
14.
15.
16.
17.
Fuel nozzle (Ref)
Ignitor
MGT terminal block
NP monopole pickup
NG monopole pickup
Anti-ice solenoid valve
Chip plug forward
18.
19.
20.
21.
22.
23.
24.
Starter-generator
Hydro mechanical unit (HMU)
Scavenge oil filter and impending bypass button
Combined engine filter assembly (CEFA)
Fuel filter
Permanent magnet alternator (PMA)
Hose assembly, combustion chamber drain (Ref)
407_MD_01_0012
Figure 1-13. Power Plant Components (Sheet 2 of 2)
31 JAN 2007—Rev. 2———1-59
BHT-407-MD-1
The bypass valves for both the fuel and
scavenge lube filters allow flow to go into a
bypass condition if excessive differential
pressure occurs across the filter elements.
The differential pressure indicators provide a
visual signal when differential pressure
across either element exceeds a
predetermined value indicating that the
element is dirty and requires replacement.
Both indicators are reset manually following
replacement.
The fuel bypass valve works in conjunction
with an impending bypass indicator. When
differential pressure across the fuel filter rises
to a pressure slightly less than the pressure
at which the bypass valve actually cracks, a
red button extends, signalling that a bypass is
about to occur. The indicator actuates at 2.1
to 2.9 PSI and the bypass valve cracks at 3.4
PSI minimum. The bypass valve seats when
the pressure drops to 3.0 PSI. The fuel filter
impending bypass indicator is located on the
lower outboard side of the manifold assembly.
The scavenge lube filter bypass indicator
v i s u a l l y s i g n a l s a n i m p e n d i n g b y pa s s
condition by extending a red button when the
differential pressure across the filter is
between 8.8 to 10.8 PSI. The lube bypass
indicator assembly is equipped with a thermal
lockout for temperatures below 43°C ±17°C
(110°F ±15°F). The lube indicator assembly is
mounted at the bottom (aft end) of the lube
bowl.
If the fuel impending bypass indicator or lube
filter bypas s in dicator exte nds , ens ure
appropriate Rolls-Royce engine maintenance
procedures are carried out prior to further
flight.
1-14. POWER PLANT IGNITION
SYSTEM
The ignition system consists of a harness, a
solid-state low tension capacitor discharge
ignition exciter, a spark igniter lead, and a
sh unte d s urfa ce-g ap spar k ig niter
(Figure 1-13).
1-60———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
The ignition system transfers energy to the
combustible fuel mixture. Energy is provided
in the form of high temperature-high
amperage arcs at the spark igniter gaps.
These arcs ignite the fuel/air mixture. The
ignition exciter is only required during the
starting cycle since the combustion process
is continuous. Once ignition takes place, the
flame in the combustion liner acts as the
ignition agent for the fuel/air mixture.
D u r i n g FA D E C m a n u a l o p e r a t i o n , t h e
automatic auto relight circuit is deactivated
and the igniter is turned on at all times at gas
producer (N G ) speeds above 55% to reduce
the possibility of engine flameout.
1-15. POWER PLANT
TE M P E R A T U R E
MEASUREMENT SYSTEM
The temperature measurement system
co nsis ts of fou r c hrome l-alume l sin gle
junction thermocouples in the gas producer
turbine outlet and an associated integral
harness. The voltages of the four
thermocouples are electrically averaged in
the assembly. The harness terminates at a
terminal block on the aft side of the horizontal
fire shield. The engine electrical harness, the
helicopter MGT indicator wiring, and the
FADEC MGT input wiring, all connect to this
terminal block (Figure 1-13).
1-16. POWER PLANT
COMPRESSOR BLEED AIR
SYSTEM
The compressor bleed air system permits
rapid engine response. The system consists
of a bleed control valve located on the front
face of the scroll and an induc er bleed
manifold which encases the slotted
compressor shroud housing.
Th e b le ed con trol v alve is o pen du ring
starting and ground idle operation, and
remains open until a predetermined pressure
ratio is ob taine d. At the predetermined
MANUFACTURER’S DATA
pressure ratio, the valve begins to modulate
from the open to the closed position.
The inducer bleed discharges air to
atmosphere at engine idle speed. At higher
power settings, flow changes from bleed to
intake air.
1-17. ENGINE OIL SYSTEM
The engine incorporates a dry sump oil
system (Figure 1-14) w ith an externally
mounted supply tank and oil cooler located
on the top aft section of the fuselage and
enclosed by the aft fairing. Oil is supplied
from the tank to gear type pressure and
scavenge pumps mounted within the engine
accessory drive gearbox. A spur gear type oil
pump assembly, consisting of one pressure
element and four scavenge elements is
mounted within.
The oil filter assembly, consisting of an oil
filter, filter bypass valve, and pressure
regulating valve, is located in the top left-hand
side of the gearbox (Figure 1-13). A check
valve, located between the filter package and
the accessory gearbox, prevents oil from
draining into the engine from the helicopter
tank when the engine is not in operation.
Additional scavenge lube filtration is provided
within the combined engine filter assembly
(CEFA) (paragraph 1-13).
Indicating type magnetic chip detectors are
installed at the bottom of the gearbox and at
the engine oil outlet connection. All engine oil
system lines and connections are internal
except the pressure and scavenge lines to the
front compressor bearing and to the bearings
in the g as p ro du c er an d po w e r tu rb in e
supports.
The system is designed to furnish adequate
lubrication, scavenging, and cooling as
needed to the bearings, splines, and gears
regardless of the helicopter attitude or
altitude. Jet lubrication is provided to all
BHT-407-MD-1
compressor, gas producer turbine, and power
turbine rotor bearings, and to the bearings
and gear meshes of the power turbine gear
train with the exception of the power output
shaft be arings. Th e pow er output shaft
bearings and all other gears and bearings are
lubricated by oil mist.
NOTE
If helicopter engine has been shut
down for more than 15 minutes,
scavenge oil could have drained into
gearbox. Dry motor run engine for 30
seconds before checking oil level. If
not a ccomplishe d, a false high
engine oil consumption rate
indication or overfilling of oil tank
could result.
The approximate capacity of the engine oil
tank is 1.5 US gallons, and the oil level is
checked by means of a sight gauge mounted
on the left side of the tank. Viewing access to
the sight gauge is provided by a cutout in the
cowling. The oil cooler is mounted on top of
the duct on the oil cooler blower.
1-18. ENGINE AUTO RELIGHT
• AUTO MODE — AUTO RELIGHT
In AUTO mode, the FADEC is capable of
detecting an engine flameout by measuring
an N G deceleration rate greater than the
predetermined flameout boundary rate. If a
fl a m e o u t i s d e t e c te d , t h e E N G I N E O U T
warning light and horn will be activated by the
FADEC ECU. Without pilot action, the auto
relight sequence is initiated, a fuel flow rate is
e s ta b li s h e d a n d t h e ig n it io n s y s t e m i s
activated. If a relight is achieved, FADEC will
control the MGT and accelerate the engine
b a c k to i ts c o m m a n d e d o p e r a ti o n . T h e
ENGINE OUT light and warning horn will turn
off after a minimum NDOT (N G acceleration
speed) or increasing MGT is established.
31 JAN 2007—Rev. 2———1-61
BHT-407-MD-1
MANUFACTURER’S DATA
OIL TANK
OIL FILTER
OIL FILTER BYPASS VALVE
CHECK VALVE
CEFA
OIL COOLER
MAGNETIC CHIP DETECTOR
PRESSURE REGULATING VALVE
OIL
PRESSURE SENSE
SCREEN
SCREEN
PRESSURE
REDUCER
OIL TANK VENT
EXTERNAL SUMP
TORQUEMETER PRESSURE
OVERBOARD BREATHER
AIR-OIL SEPARATOR
MAGNETIC CHIP DETECTOR
SUPPLY AND
BYPASS OIL
PRESSURE OIL
SCAVENGE OIL
TORQUE
MODULATED
PRESSURE
SCAVENGE
RETURN
VENT TO
ATMOSPHERE
TYPICAL
407_MD_01_0013
Figure 1-14. Engine Oil System
1-62———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
The automatic auto relight sequence will
initiate from detection of flameout until the NG
speed decays to 50%. Once the N G decays
below 50%, the FADEC will no longer attempt
to relight the engine. In the event of an
unsuccessful relight refer to BHT-407-FM-1,
Section 3.
• MANUAL MODE
In MANUAL mode, the FADEC controlled auto
relight circuit is disabled. Because of this, the
ignition system has been designed to operate
continuously in MANUAL mode, at engine gas
producer (N G ) speeds of 55% or greater to
reduce the possibility of flameout.
When the FADEC system is in MANUAL mode,
the NG gauge operates as a trigger device for
the engine out horn and light when NG drops
below 55%. In the event of a power loss to the
N G indicator while operating in MANUAL
mode, the failure mode will provide
continuous ignition regardless of NG speed,
but the ENGINE OUT light and horn will not be
activated when NG drops below 55%.
• AUTO MODE — PILOT ASSISTED IN-FLIGHT
RESTART
In addition to the above mentioned Engine
Relight features, the FADEC system also
incorporates specific relight logic, for engine
out conditions, when the NG speed is between
9.5% and 50%. Pilot action is required to
initiate an in-flight restart at NG speeds below
50%.
When appropriate BHT-407-FM-1 procedures
are followed, the in-flight restart logic will
introduce fuel scheduling based on the
ex is ti ng N G s p ee d . S ho u ld a r eli gh t b e
achieved, the FADEC will accelerate the
engine to an idle speed of 63% N G. As the
priority of the in-flight relight logic is to help
achieve an engine start in an emergency
condition, the hot start abort function is
disabled.
Therefore, to help reduce the possibility of an
over temperature condition from occurring,
BHT-407-MD-1
BHT-407-FM-1 procedures require that the
throttle be initially positioned to the closed
position and the start switch be positioned to
STA RT. Once the starter is assisting to
maintain or increa se the N G spee d, the
throttle can be positioned to IDLE and the
FADEC will introduce fuel scheduling. Ignition
will be provided in conjunction with activation
of the starter.
As the in-flight restart logic is designed for NG
speeds between 9.5 and 50% NG, if an in-flight
restart is initiated below 9.5% NG, normal start
logic will be used to introduce fuel based on
5.202 or 5.356 software (paragraph 1-12-D).
S h ou ld a r e lig h t oc c u r, th e FA D E C w ill
accelerate the engine to idle. In addition, hot
start abort logic will be enabled for starts
initiated at NG speeds below 9.5%.
1-19. AUTO RELIGHT CAUTION
LIGHT
The AUTO RELIGHT light will be ON when the
ignition system is activated.
1-19-A. START IN AUTO MODE
During start in the AUTO mode, the ignition
system is activated via the engine igniter
circuit breaker and the start circuit until the
NG speed reaches 50 ±1%. Above this speed,
the FADEC carries out an auto relight test and
continues activation of the ignition system
until a gas producer (NG) speed of 60 ±1%, at
which time the AUTO RELIGHT light will go
OFF.
1-19-B. ENGINE OUT IN AUTO MODE
In the event of an engine out condition during
normal engine operations with the FADEC in
AUTO mode, the ignition system is activated
by the FADEC. This will occur when an engine
out condition is detected and the engine is
restarted or until the gas producer (NG) speed
decays to 50%. The ENGINE OUT warning
l ig h t a n d h o r n a n d t h e A U TO R E L IG H T
advisory light will be on at all times under
these conditions.
31 JAN 2007—Rev. 2———1-63
BHT-407-MD-1
1-19-C. START AND CONTINUOUS
OPERATION IN MANUAL MODE
During a start in MANUAL mode, the ignition
system will be activated when the starter
switch is positioned and held in START.
F o l l o w i n g t h e e n g i n e s ta r t a n d d u r i n g
continuous operation at gas producer (N G )
speeds above 55%, the ignition system will
operate continuously. The AUTO RELIGHT
light will be on at all times under these
conditions.
1-20. ENGINE ANTI ICE SWITCH
The ENG ANTI ICE switch, located in the
ov e rh e ad co n s ol e, c o nt ro ls t he e ng in e
anti-ice bleed air solenoid valve. The engine
anti-ice system will be activated when the
ENG ANTI ICE switch is positioned to ENG
ANTI ICE. This de-energizes the engine
anti-ice bleed air solenoid valve allowing hot
diffuser scroll air to flow from the engine
anti-icing air valve to the engine compressor
front support guide vanes and prevent the
fo rmation of ice. In the ev ent of a total
electrical system failure, the anti-ice system
will fail safe to ON and provide continuous
anti-icing.
When the ENG ANTI ICE switch is positioned
to OFF, bus voltage is provided to the engine
anti-ice bleed air solenoid valve from the
engine anti-ice circuit breaker. This energizes
the engine anti-ice bleed air solenoid valve
and prevents the flow of hot air from the
e n g i n e a n t i - i c e a i r v a lv e t o t h e e n g i n e
compressor front support guide vanes.
On helicopters incorporating an ENGINE ANTI
ICE caution panel annunciator, an engine
anti-ice pressure switch is used to control
activation of the annunciator. Engine anti-ice
pressure switch activation occurs on
increasing pressure at 5.5 ±.5 PSI (37.9 ±3.4
kPa), which allows the ENGINE ANTI ICE
annunciator to illuminate. Engine anti-ice
pressure switch deactivation occurs on
decreasing pressure prior to 3.0 PSI (20.68
1-64———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
kPa), which turns OFF the ENGINE ANTI ICE
annunciator.
1-21. ENGINE INDICATORS
Engine indicators include a TORQUE, N G,
MGT, N R /N P, engine oil temperature and
pressure gauge, ENGINE OUT warning light,
and horn, ENGINE OVSPD light, and ENGINE
CHIP light.
1-21-A. TORQUE GAUGE
The engine torque meter pressure sensing
port outputs a specific oil pressure for a
specific engine torque. This oil pressure is
f e d t o b o t h t h e FA D E C a n d a i r f r a m e
t r a n s d u c e r s . T h e FA D E C t r a n s d u c e r
information provides input to the ECU for the
operation of the FADEC system. The airframe
transducer provides a signal to the torque
gauge.
1-21-A-1.
TORQUE GAUGE
MARKING
—
RANGE
The 100% maximum on the torque gauge
r e p r e s e n ts 5 6 0 f o o t -p o u n d s ( 6 7 4 s h a ft
horsepower at 100% NP) at the engine output
shaft.
This output is divided on the helicopter
between the tail rotor drive system, the main
rotor drive system, accessory requirements
and gearbox losses.
The 93.5% on the torque gauge represents
524 foot-pounds (630 shaft horsepower at
100% NP). This is the maximum limit allowed
on a continuous basis on the engine gearbox.
T h e 5 m i n ut e lim it b e tw ee n 9 3 .5 % (5 2 4
foot-pounds) and 100% (560 foot-pounds)
represents a 5-minute limit on the engine
gearbox in this range of power.
Any torque exceedance value recorded by the
indicator represents the straight torque seen
by both the engine gearbox and the main
transmission. In the event of a recorded
torque exceedance, reference to both the
BHT-407-MM and Rolls Royce 250-C47B
MANUFACTURER’S DATA
Operation and Maintenance Manual is
required to ensure appropriate inspections
are carried out.
NOTE
Litton instrument torque does not
equal FADEC ECU torque for the
same input oil pressure. This is due
to the fact that the Litton torque
instrument and the FADEC ECU use
separate torque transducers to arrive
at their respective indicated torques
(i.e. engine limit = 590 foot-pounds/
710 HP and transmission limit = 560
foot-pounds/674 HP). To convert
from Litton instrument% torque to
FADEC ECU% torque, divide Litton
instrument% torque by 1.0535. To
convert FADEC ECU% torque to
foot-pounds, multiply FADEC ECU%
tor qu e by 5 .9 . T o co n ve rt f rom
FADEC ECU% torque to Litton
instrument torque, multiply FADEC
ECU% torque by 1.0535.
1-21-A-2.
ENGINE TORQUE EXCEEDANCE
The torque gauge microprocessor is pre
programmed with specific torque values
(Table 1-4).
BHT-407-MD-1
There is a torque limitation of 5 minutes
between 93.5% and 100%. To give the pilot
notice that he is approaching the end of the 5
minute limit, the Trend ARC of indicator will
flash and CHECK INSTR light will come on
when 30 seconds remain in the 5 minute time
period.
There is a redline at 100% representing the
maximum torque allowed. When the pilot
exceeds 100%, the indicator Trend ARC will
immediately begin to flash and CHECK INSTR
light will illuminate.
There is an engine gearbox limit at 105.4%.
When the pilot exceeds 100%, the Trend ARC
of indicator will already be flashing. However,
after 10 seconds at or above 105.4%, the
indicator will indicate that an exceedance has
occurred and will begin to record the
exceedance in NVM. At this time, the indicator
Trend ARC will stop flashing.
There is a transmission inspection limit at
110%. When the pilot exceeds 100%, the
Tr e n d A R C o f i n d ic a t o r w il l a lr e a d y b e
flashing. Any exceedance above 110% will
immediately be recorded in NVM. The
indicator Trend ARC will stop flashing after 10
seconds above 105.4% as described above.
Table 1-4: Engine Torque Exceedance Monitoring
Torque%
Indication
Exceedance
93.6 to 100%
Trend ARC begins flashing after 4.5
minutes.
Not recorded.
100.1 to 105.3%
Trend ARC begins flashing immediately
and flashes continuously as long as in
this range.
Not recorded.
105.4 to 110%
Trend ARC begins flashing immediately.
Flashes for 10 seconds, then stops
flashing.
Recorded after 10 seconds.
Above 110.1%
Trend ARC flashes prior to reaching
110.1%, as described above, but stops
flashing at 110.1% or above.
Recorded immediately.
31 JAN 2007—Rev. 2———1-65
BHT-407-MD-1
MANUFACTURER’S DATA
1-21-B. GAS PRODUCER GAUGE (NG)
The gas producer (NG) gauge displays engine
gas producer speed in percent of rated RPM
(Table 1-5).
An N G Speed Pickup mounted in the engine
gearbox provides two separate signals to the
ECU. One signal is the primary driver of the
NG gauge and is also a secondary signal for
the ECU. The second signal is the primary
input to the ECU. The primary signal to the
gauge will be passed through the ECU, even
when the ECU is not powered.
When the FADEC system is in MANUAL mode,
the NG gauge operates as a trigger device for
the ENGINE OUT horn and light when N G
drops below 55%. In the event of a power loss
to the NG indicator while operating in
MANUAL mode, the failure mode will provide
continuous ignition regardless of NG speed,
but the ENGINE OUT light and horn will not be
activated when NG drops below 55%.
Table 1-5: GAS PRODUCER GAUGE (NG) Exceedance Monitoring
N G%
Indication
Exceedance
105.1 to 106%
Trend ARC starts flashing immediately.
Stops flashing after 10 seconds.
Recorded after 10 seconds.
Above 106.1%
Trend ARC begins flashing prior to
reaching 106.1%, as described above.
Stops flashing at 106.1% and above.
Recorded immediately.
1-21-C. DUAL TACH GAUGE
The Dual Tachometer has two pointers: one
that displays Main Rotor RPM on the outer
scale and one that displays engine NP on the
inner scale. All scales are in percent of rated
RPM. If Quiet Cruise Mode kit is installed,
refer to the BHT-407-FMS-25.
The NP side of the instrument is powered by
the 28 VDC bus throu gh its own circuit
breaker.
An NP Speed Pickup mounted on the engine
provides two separate signals to the ECU.
One signal is the primary driver of the N P
gauge and also a secondary signal for the
ECU. The second signal is the primary input
to the ECU. The primary signal to the gauge
will be passed through the ECU, even when
the ECU is not powered.
The NR side of the gauge is powered by the 28
VDC bus through its own circuit breaker.
1-66———Rev. 2—31 JAN 2007
An NR speed pickup mounted on the
transmission lower case provides three
separate identical signal outputs of rotor
RPM. One signal is sent to the NR circuit of
the dual tachometer gauge. One signal is sent
to the ECU and the other is sent to the Main
Rotor RPM sensor switch.
1-21-D. MEASURED GAS TEMPERATURE
GAUGE
The measured gas temperature (MGT) gauge
dis play s e ng ine ga s te mper ature o f air
between gas producer turbine and power
turbine in degrees Celsius.
The MGT gauge has one set of exceedance
levels for starting and another for normal
operation (Table 1-6 through Table 1-9). The
indicator uses the start exceedance levels
when the FADEC/START relay is engaged
during normal starts.
MANUFACTURER’S DATA
BHT-407-MD-1
Table 1-6: MGT Exceedance Monitoring — During Start (P/N 407-375-001-101/103)
MGT°C
Indication
Exceedance
826 to 926°C
Trend ARC begins flashing after 6 seconds. Recorded after 10 seconds.
Stops flashing after 10 seconds.
926.1 to 927.9°C
Trend ARC begins flashing prior to
reaching 926.1°C, as described above but
stops flashing at 927.9°C and above.
Recorded after 1 second.
Above 928°C
Trend ARC begins flashing prior to
reaching 928°C, as described above but
stops flashing at 928°C or above.
Recorded immediately.
Table 1-7: MGT Exceedance Monitoring — During Normal Operation (P/N 407-375-001-101/103)
MGT°C
Indication
Exceedance
727.1 to 779°C
Trend ARC begins flashing after 4.5
minutes. Stops flashing after 5 minutes.
Recorded after 5 minutes.
779.1 to 826°C
Trend ARC begins flashing immediately
and stops flashing after 12 seconds.
Recorded after 12 seconds.
Above 826.1°C
Trend ARC begins flashing prior to
reaching 826.1°C, as described above but
stops flashing at 826.1°C or above.
Recorded immediately.
Table 1-8: MGT Exceedance Monitoring — During Start (P/N 407-375-001-105 and Subsequent)
MGT°C
Indication
Exceedance
843 to 926°C
Trend ARC begins flashing after 6 seconds.
Stops flashing after 10 seconds.
Recorded after 10 seconds.
926.1 to 927.9°C
Trend ARC begins flashing prior to
reaching 926.1°C, as described above, but
stops flashing at 927.9°C and above.
Recorded after 1 second.
Above 928°C
Trend ARC begins flashing prior to
reaching 928°C as described above, but
stops flashing at 928°C or above.
Recorded immediately.
Table 1-9: MGT Exceedance Monitoring — During Normal Operation
(P/N 407-375-001-105 and Subsequent)
MGT°C
Indication
Exceedance
727.1 to 779°C
Trend ARC begins flashing after 4.5
minutes. Stops flashing after 5 minutes.
Recorded after 5 minutes.
779.1 to 905°C
Trend ARC begins flashing immediately
and stops flashing after 12 seconds.
Recorded after 12 seconds.
31 JAN 2007—Rev. 2———1-67
BHT-407-MD-1
MANUFACTURER’S DATA
Table 1-9: MGT Exceedance Monitoring — During Normal Operation
(P/N 407-375-001-105 and Subsequent) (Cont)
MGT°C
Indication
Exceedance
Above 905°C
Trend ARC begins flashing prior to
reaching 905°C, as described above, but
stops flashing at 905°C or above.
Recorded immediately.
1-21-E. ENGINE OIL TEMPERATURE/
PRESSURE GAUGE
The engine oil temperature and pressure
gauge is a dual instrument that
simultaneously displays oil temperature in
degrees Celsius on the right side display and
oil pressure in PSI on the left side display.
Each side of the indicator is powered by its
o w n c i r c u i t b r e a k e r. T h e e n g i n e o i l
temperature input signal is provided by a
thermobulb installed on the engine oil tank.
The engine oil pressure is provided by a
transducer mounted on the forward engine
fire wall.
1-22. ENGINE OUT WARNING LIGHT
AND HORN
The ENGINE OUT light and (pulsing) warning
horn circuit will activate when the FADEC
detects an engine flameout (by sensing N G
deceleration) or when gas producer (N G )
speed is 55 ±1% or less.
1-22-A. FADEC IN AUTO MODE
If the FADEC detects an engine flameout (NG
deceleration) or NG speed of 55 ±1% or less,
the ENGINE OUT light and (pulsing) warning
horn circuit will activate. In addition, the
FADEC will automatically initiate an auto
relight sequence immediately upon detection
of a flameout (NG deceleration).
1-22-B. FADEC IN MANUAL MODE
In the MANUAL mode, the FADEC engine out
detection circuit is deactivated and activation
o f th e E N G I N E O U T lig h t a n d (p u ls i ng )
warning horn is controlled by the gas
producer NG gauge (NG less than 55 ±1%).
1-68———Rev. 2—31 JAN 2007
Engine out (flameout) protection in the
MANUAL mode is provided by continuous
activation of the ignition system at NG speeds
of 55% or greater.
The engine out (pulsing) warning horn can be
muted in either the AUTO or MANUAL mode
by pressing the HORN MUTE switch.
1-23. ENGINE OVERSPEED
WARNING LIGHT
The ENGINE OVSPD light will illuminate if the
FADEC detects a NG overspeed of 110 ±1% or
a NP overspeed of 118.5 ±1% (5.202 and 5.356
FADEC software). Illumination occurs when
the overspeed solenoid valve is activated
within the Hydro Mechanical Unit (HMU). The
ENGINE OVSPD light will also illuminate when
NP versus TORQUE (5.202 FADEC software) is
above the maximum continuous limit (102.1%
NP at 100% torque to 108.6% NP at 0% torque)
or when NP is above the maximum continuous
limit of 102.1% (5.356 FADEC software). The
ENGINE OVSPD light is operational with the
FADEC in AUTO or MANUAL mode.
If the ENGINE OVSPD light is activated during
engine operation, due to an exceedance, it
will be recorded by the ECU and the pilot will
be provided with a maintenance advisory on
shutdown in the form of a FADEC DEGRADED
lig ht. T h e FA D E C D E GR A D E D lig ht w ill
illuminate when NG speed decays below 9.5%.
If the pilot fails to recognize illumination of
the FADEC DEGRADED light on shutdown, it
will be illuminated the next time electrical
power is applied following the FADEC system
self test. When the FADEC DEGRADED light is
illuminated as a maintenance advisory,
maintenance action is required prior to further
flight. Peak values of exceedances are located
MANUFACTURER’S DATA
on the E ngine History Da ta page of the
EMC-35A Maintenance Terminal.
The ENGINE OVSPD light will also illuminate
momentarily during the overspeed shut down
test when the overspeed solenoid valve is
activated within the HMU (BHT-407-FM-1).
1-24. ENGINE CHIP CAUTION LIGHT
The ENGINE CHIP light will illuminate if
metallic particles in the oil accumulate on
either the engine sump or scavenge chip
detectors. The engine sump chip detector is
located on the lower left side of the engine
ge a rbo x a nd the en gi ne s ca v en ge ch ip
detector is located on the forward right-hand
side of the engine gearbox. Both chip
detectors are a quick disconnect design.
1-25. FUEL SYSTEM DESCRIPTION
The fuel system (Figure 1-15) consists of two
crash resistant, bladder type fuel cells. The
forward fuel cell is located underneath and
between the aft facing passenger seats. The
aft fuel cell is located underneath and behind
the aft passenger seats.
Both fuel cells are serviced through the filler
port located on the right side of the helicopter.
Approximately 28.4 US gallons (193.1 lb) of
fuel will accumulate in the aft main fuel cell,
prior to the forward cell being filled through
the gravity feed stand pipe. The gravity feed
stand pipe connects the aft fuel cell to the
forward tank. As the aft fuel cell fills beyond
the level of the top of the stand pipe, the
forward tank will be completely filled (forward
tank fuel capacity 37.6 US gallons (256.0 lb).
The aft tank will then be filled to the level of
the filler port (usable fuel capacity 127.8 US
gallons (869.0 lb).
If the auxiliary tank is installed, it will be filled
at the same time as the aft tank through two
openings located on the lower aft wall of the
BHT-407-MD-1
main fuel cell, which are connected to the
b o t t o m o f t h e a u x i l i a r y ta n k ( r e f e r t o
BHT-407-FMS-6 for information on this kit).
Fuel from the forward tank is transferred to
the main tank by two transfer pumps mounted
on a sump plate assembly located on the
bottom of the forward fuel cell. Fuel from the
main fuel cell will be supplied to the engine
through two boost pumps located at the base
of the main fuel cell on a sump plate. The fuel
fr o m th e t w o b o o s t p u m ps jo i n s in t o a
common fuel line that passes through a fuel
shutoff valve, then through an airframe
mounted fuel filter before reaching the engine
driven pump on the HMU. A solenoid sump
drain valve is installed on the sump plate
assemblies of both the forward and aft fuel
tanks. The drain valves are activated by two
switches located on the right hand side of the
lo w e r a ft f us e la g e. T h e se s w itc h es a re
deactivated when the fuel valve switch is ON
to prevent inadvertent activation during flight.
1-25-A. FUEL SYSTEM OPERATION
With power applied to the helicopter and the
Fuel XFR/Boost Right and Fuel XFR/Boost
Left circuit breaker switches ON, the two
transfer pumps and the two boost pumps are
operating. Refer to Figure 1-16 for the fuel
transfer system schematic.
Each transfer pump sends fuel through a
one-way check valve located at the outlet of
each pump and a common tee fitting through
a line which transfers the fuel into the aft tank.
Each check valve will assure that if either of
the transfer pumps becomes inoperative fuel
will be pumped to the aft tank and not through
the inoperative pump back into the fwd fuel
tank.
A modified tee fitting incorporating an orifice
allows for a specific amount of fuel to bleed
back into the forward tank to reduce the
transfer rate.
30 APR 2008—Rev. 4———1-69
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL VALVE SWITCH
FUEL
PRESS
FUEL VALVE
(AMBER)
FIREWALL
FUEL FILTER
(AMBER)
AIRFRAME
FUEL FILTER
ENGINE
MANUAL
DRAIN VALVE
VENT LINE
DRAIN
HOSE
SHUT OFF VALVE WITH
THERMAL RELIEF
PRESSURE
TRANSDUCER
FUEL
QTY
FILLER CAP
LBS
ENGINE FEED
LINE
FWD FUEL
CELL
FUEL SIGNAL
CONDITIONER
AUX FUEL TANK
KIT (REF)
OVERBOARD
VENT
AFT FUEL
CELL
QTY PROBE
(3 PLACES)
CHECK VALVE
WITH THERMAL
RELIEF
TRANSFER
BLEED
TRANSFER
PUMP DUAL
CHECK VALVE
WITH THERMAL
RELIEF
TRANSFER
LINE
BOOST PUMP DUAL
ELECTRIC
SUMP DRAIN
PRESSURE
SWITCH
LOW LEVEL
SWITCH
INTERCONNECT LINE
R/FUEL XFR
(AMBER)
L/FUEL XFR
(AMBER)
ELECTRIC SUMP
DRAIN
L/FUEL BOOST
(AMBER)
LOW LEVEL SWITCH
R/FUEL BOOST
(AMBER)
FUEL LOW
(AMBER)
TYPICAL
407_MD_01_0014
Figure 1-15. Fuel System
1-70———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
RIGHT XFR
PUMP RELAY
FUEL XFR/
BOOST RIGHT
28 VDC BUS
RIGHT XFR
PUMP
10
FUEL QTY IND
3
PUMP OUTPUT
PRESSURE
FUEL SIGNAL
CONDITIONER
TIME
DELAY
CAUTION
WARNING PANEL
POWER IN
L
R/FUEL XFR
D
L/FUEL XFR
RIGHT FUEL
PRESSURE SWITCH
TIME
DELAY
POWER IN
LEFT FUEL
PRESSURE SWITCH
FUEL
PROBE
FORWARD FUEL TANK
LEFT XFR
PUMP RELAY
FLOAT
PUMP OUTPUT
PRESSURE
FUEL XFR/
BOOST LEFT
LOW LEVEL
DETECTOR
10
LEFT XFR
PUMP
BATTERY SWITCH
BATTERY POWER
OFF
4
5
28 VDC POWER
S/N 53000 THROUGH 53174
6
ON
TYPICAL
407_MD_01_0015
Figure 1-16. Fuel Transfer System Schematic (Sheet 1 of 2)
31 JAN 2007—Rev. 2———1-71
BHT-407-MD-1
MANUFACTURER’S DATA
28 VDC BUS
RIGHT XFR
PUMP
RIGHT XFR
PUMP RELAY
FUEL XFR/
BOOST RIGHT
10
FUEL QTY IND
3
PUMP OUTPUT
PRESSURE
FUEL SIGNAL
CONDITIONER
TIME
DELAY
CAUTION
WARNING PANEL
POWER IN
R/FUEL XFR
D
L/FUEL XFR
RIGHT FUEL
PRESSURE SWITCH
GROUND CIRCUIT
FOR XFR LIGHTS
BASED ON
ACTIVATION TABLE
TIME
DELAY
L
POWER IN
LEFT FUEL
PRESSURE SWITCH
FUEL
PROBE
FORWARD FUEL TANK
LEFT XFR
PUMP RELAY
FLOAT
PUMP OUTPUT
PRESSURE
FUEL XFR/
BOOST LEFT
LOW LEVEL
DETECTOR
10
LEFT XFR
PUMP
BATTERY SWITCH
BATTERY POWER
OFF
4
5
28 VDC POWER
S/N 53175 AND SUBSEQUENT
6
ON
TYPICAL
407_MD_01_0016
Figure 1-16. Fuel Transfer System Schematic (Sheet 2 of 2)
1-72———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
Eac h bo ost pu mp sen ds fue l th rough a
one-way check valve located at the outlet of
each pump and a common tee fitting. It then
flows through a common line to a fitting
located at the top right side of the main fuel
cell. The line leaves the fuel cell and passes
through the fuel shut off valve. A pressure
transducer (located between the main fuel cell
fitting and the fuel shut off valve) supplies the
electrical signal to the instrument panel
mounted fuel pressure gauge.
During normal operation, fuel will first be
used from the aft fuel cell until its level is
equivalent to the top of the forward tank. At
this time, the aft cell and forward cell will be
used equally down to the level of the top of
the gravity feed stand pipe. At this point, all
fuel will be used from the forward tank.
Finally, after the forward tank is empty, the
remaining fuel in the aft tank will be used.
Refer to paragraph 1-25-P-2 for operational
information of the L/FUEL XFR and R/FUEL
X F R L I G H T S a n d pa r a g r a p h 1 - 2 5 -R f o r
information on the FUEL LOW LIGHT.
1-25-B. LEFT FUEL BOOST/XFR
ALTERNATE ELECTRICAL
CIRCUIT
In the event that a short circuit or battery hot
condition occurs in the helicopter and all DC
bus power is shut off, it is desirable to
maintain the operation of one transfer pump
and one boost pump.
The BATT switch configures the DC power
feed to the left fuel boost pump and the left
fuel transfer pump between the DC bus and
the helicopter battery (Figure 1-17).
In the event the battery switch is positioned to
OFF during helicopter operations, an
alternate circuit is provided to allow operation
of the left fuel transfer and left fuel boost
pumps. During this condition, with the fuel
valve switch positioned to ON, battery voltage
BHT-407-MD-1
is supplied through the FUEL BOOST/XFR
backup circuit breaker, the fuel valve switch,
the battery switch, and the left fuel XFR/boost
circuit breaker switch to the left fuel transfer
and left fuel boost pumps.
1-25-C. FUEL SYSTEM CONTROLS
Fuel system controls are located on the
instrument panel, overhead console, and aft
lower right fuselage. The controls consist of a
FUEL VALVE switch, a FUEL BOOST/XFR
LEFT switch, a FUEL BOOST/XFR RIGHT
switch, and two FUEL CELL DRAIN switches.
1-25-D. FUEL VALVE SWITCH
The FUEL VALVE switch is located on the
lower right side of the instrument panel. It is a
guarded two position toggle switch that
provides a means of shutting off the flow of
fuel to the engine.
The fuel shutoff valve is a motorized gate
valve which must be driven to either the open
or closed position. When activated by the
FUEL VALVE switch, the shutoff valve motor
will rotate in the appropriate direction to
OPEN or CLOSE the gate valve. The FUEL
VALVE light will momentarily illuminate
during valve transit.
1-25-E. LEFT AND RIGHT FUEL BOOST/
XFR SWITCH
The left and right FUEL BOOST/XFR switches
(Figure 1-16) are a circuit breaker toggle
design and located on the overhead console.
Each switch controls the operation of its
respective boost pump located in the main
fuel tank and transfer pump in the forward
fuel tank. In addition to the left and right
FUEL/XFR switches, separate electrical
circuits within the fuel signal conditioner are
used to control the operation of the left and
right fuel transfer pumps as the forward fuel
tank is emptied.
31 JAN 2007—Rev. 2———1-73
BHT-407-MD-1
The LEFT FUEL/XFR switch is outlined by a
yellow border (Figure 1-5) to identify that it
has an alternate circuit (Figure 1-17). In the
event the BATT switch is positioned to OFF
during helicopter operations, an alternate
circuit is also provided to allow operation of
the left fuel transfer and left fuel boost pumps.
During this condition, with the FUEL VALVE
switch positioned to ON, battery voltage is
supplied through the FUEL BOOST/XFR
BACKUP circuit breaker (located on the left
side of the instrument pedestal above the chin
bubble), the FUEL VALVE switch, the BATT
SWITCH, and the LEFT FUEL BOOST/XFR
circuit breaker switch to the left fuel transfer
and left fuel boost pumps. Therefore, in this
situation the LEFT BOOST pump will continue
to run. The LEFT XFR pump will continue to
run as long as the fuel signal conditioner
determines that there is fuel in the forward
tank.
1-25-F. FUEL CELL DRAIN SWITCHES
The FUEL CELL DRAIN switches provide a
means of draining fuel from both the forward
and aft fuel cells individually. The switches
are a momentary push button design made up
of an outer environmental seal and an inner
switch assembly. They are mounted side by
side on the lower right side of the aft fuselage,
below and aft of the fuel filler port.
The FUEL VALVE switch must be in the OFF
position to operate either of the fuel cell drain
valves. This prevents inadvertent operation of
the fuel cell drains during flight.
With helicopter electrical power provided and
the FUEL VALVE switch positioned to OFF,
pressing the FUEL CELL DRAIN switches will
OPEN the respective drain valves allowing
fuel to be drained from the forward and main
fuel cells.
1-74———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
1-25-G. FUEL SYSTEM INDICATORS
Fuel system indicators consist of a fuel
quantity indicator, fuel pressure indicator,
FUEL VALVE light, L/FUEL BOOST light,
R/FUEL BOOST light, L/FUEL XFR light, R/
FUEL XFR light, FUEL FILTER light, and a
FUEL LOW light.
1-25-H. FUEL SYSTEM CAPACITY
Fo r fuel s yste m c apac ity refer to the
BHT-407-FM-1, Section 5.
1-25-I.
FUEL QUANTITY GAUGING
SYSTEM (FQGS)
The fuel quantity gauging system (FQGS)
measures the quantity of fuel in the two main
fuel tanks. The FQGS also measures the
quantity of fuel in the auxiliary fuel tank when
it is installed.
Th e fue l qu antity is mea sure d b y th ree
capacitance type probes in the fuel tanks. The
signals from the probes are used by the fuel
signal conditioner to calculate the fuel weight.
The signal conditioner provides a signal to
the fuel quantity gauge to display the
calculated fuel weight.
1-25-J. FUEL QUANTITY SIGNAL
CONDITIONER
The signal conditioner is a separate unit that
is located on the aft electrical equipment shelf
next to the DC controller (Voltage Regulator).
The electronic interface circuits for the fuel
low level detection system are located along
with the fuel gauging system signal
conditioner circuits in the same physical unit.
Both systems are physically and electrically
separate within the unit. There are three LEDs
located on the aft side of the signal
conditioner which are used to display the
FQGS status for purposes of troubleshooting.
MANUFACTURER’S DATA
BHT-407-MD-1
BATTERY
FUEL VALVE SWITCH
OFF
1
+
BATTERY RELAY
FUEL BOOST/
XFR BACKUP
A2
15
2
3
4
5
BATTERY SWITCH
4
6
ON
OFF
5
28 VDC BUS
6
ON
FUEL XFR/
BOOST LEFT
10
LEFT
BOOST PUMP
LEFT
TRANSFER PUMP
LEFT XFR
PUMP RELAY
FUEL SIGNAL
CONDITIONER
TYPICAL
407_MD_01_0017
Figure 1-17. Left Fuel Boost/XFR Alternate Circuit Schematic
31 JAN 2007—Rev. 2———1-75
BHT-407-MD-1
MANUFACTURER’S DATA
1-25-K. SIGNAL CONDITIONER
BUILT-IN-TEST (BIT)
has been corrected unless the power is
turned OFF and ON to the signal conditioner.
1-25-K-1.
In addition, the failures detected will be
displayed on three LEDs on the back of the
signal conditioner (refer to the BHT-407-MM
for fault explanation).
POWER-UP BIT
The signal conditioner receives power from
the 28 VDC bus through the FUEL QTY INSTR
circuit breaker. The signal conditioner carries
out a power up BIT when the unit is first
provided power. The power-up BIT must be
completed before the signal conditioner can
take any readings of fuel quantity. The signal
conditioner should complete a power-up BIT
check within approximately 4 seconds after
application of power. There is no connection
between the BIT feature of the signal
conditioner and the BIT performed by the fuel
quantity gauge.
If a failure is detected during the power-up BIT
or if an error is found in the probe signal
received or the power source for the probe
input, the signal conditioner will blank the fuel
quantity gauge.
If errors have been detected and the gauge
has been blanked, the signal conditioner will
not turn the gauge back on even if the error
has been corrected unless the power is
turned OFF and ON to the signal conditioner.
In addition, the failures detected will be
displayed on three LEDs on the back of the
signal conditioner (refer to the BHT-407-MM
for fault explanation).
1-25-K-2.
CONTINUOUS BIT
1-25-L. FUEL QUANTITY CALCULATION
A microprocessor in the signal conditioner
uses the information provided by the three
fuel probes to compute the weight of the fuel
in the following steps.
1.
Uses the Main Tank Forward Fuel
Probe (Probe No. 2) input signal for
density correction if it is totally
immersed in fuel or if not, uses a
default density (default density is
6.6594 pounds/gallons).
2.
Calculates the height of the fuel
indicated for each of the three probes
corrected for fuel density by using
the value from step 1.
3.
The calculated height on each probe
is used to look up a volume of fuel in
gallons in a table contained in the
NVM of the signal conditioner.
4.
The weight of the fuel is then
calculated by multiplying the volume
by the density. A calculation is done
for all three probes to compute the
total system weight of fuel. A
calculation is also done for only the
Forward Tank Fuel Probe (Probe
No. 3).
The signal conditioner carries out a
continuous BIT whenever it is powered. If a
failure is detected during the continuous BIT
or if an error is found in the probe signal
received or the power source for the probe
input, the signal conditioner will blank the
gauge display.
To ta l s y s t e m w e i g h t i n f o r m a t i o n i s
transmitted from the signal conditioner to the
fuel quantity gauge. When the forward fuel
quantity switch is pushed, the weight for the
forward tank only is transmitted.
If errors have been detected and the gauge
has been blanked, the signal conditioner will
not turn the gauge back on even if the error
The usable fuel weight (in pounds) is
calculated by the fuel signal conditioner and
is displayed by the gauge.
1-76———Rev. 2—31 JAN 2007
1-25-M. FUEL QUANTITY GAUGE
MANUFACTURER’S DATA
BHT-407-MD-1
1-25-N. FUEL QUANTITY BUTTON
The fuel quantity gauge normally indicates
the total usable fuel in both fuel tanks and
auxiliary tank (if installed). Pushing the FUEL
QTY FWD TANK button will make the fuel
quantity gauge display the fuel in the forward
tank only.
1-25-O. FUEL PRESSURE/AMMETER
GAUGE
The fuel pressure/ammeter gauge is a dual
display gauge. The left side of the instrument
displays the pressure output from the two fuel
boost pumps in pounds per square inch (PSI).
The fuel pressure side of the instrument is
powered by its own circuit breaker. The
instrument receives its input signal from a
transducer mounted between the fuel boost
pu mps an d the fu el shu t off va lve. Th e
minimum pressure limit is set to ensure that
sufficient fuel pressure will be supplied to the
input of the engine driven fuel pump.
1-25-P. FUEL VALVE LIGHT
The FUEL VALVE light will illuminate when the
fuel shut off valve is in transit or has stopped
somewhere between the full OPEN or full
CLOSED position.
1-25-P-1.
L/FUEL
LIGHTS
AND
R/FUEL
BOOST
The L/FUEL BOOST and R/FUEL BOOST
lights will illuminate when their respective
fuel pressure switch senses a decreasing
boost pump output pressure of 1.5 ±0.5 PSI.
When boost pump pressure is increasing, the
lights will extinguish prior to the pressure
passing through 5 PSI.
1-25-P-2.
1-25-P-2-A.
L/FUEL XFR AND R/FUEL XFR
LIGHTS
S/N 53000 THROUGH 53174
The L/FUEL XFR and R/FUEL XFR lights
(Figure 1-16, Sheet 1) will illuminate when
their respective fuel pressure switch senses a
decreasing boost pump output pressure of
1.5 ±0.5 PSI. When boost pump pressure is
increasing the lights will extinguish prior to
the pressure passing through 5 PSI.
Fuel must be present in the forward fuel cell
for the annunciator circuits to operate since
they are both controlled by the fuel signal
conditioner. When the forward fuel cell is
nearing depletion the signal conditioner uses
time delays to control the transfer pumps and
lights. The time delay allows continued
operation of the transfer pumps to ensure all
of the fuel in the forward tank is transferred to
the main fuel tank. The forward fuel cell will be
empty when approximately 193.1 pounds of
total fuel is indicated.
The fuel signal conditioner time delay which
controls the L/FUEL XFR pump circuit is
provided by the FUEL XFR/BOOST LEFT
circuit breaker. The fuel signal conditioner
time delay which controls the R/FUEL XFR
pump circuit is provided by the FUEL XFR/
BOOST RIGHT circuit breaker.
T he L /F U EL X FR an d R /F U E L X FR lig ht
circuits will remain operational during the
time delay and will illuminate in the event
transfer pump output pressure drops below
1.5 ±0.5 PSI.
Once the time delay periods are ended, the
transfer pumps and the L/FUEL XFR and R/
FUEL XFR light circuits are deactivated. The
transfer pump and light circuits will stay
inoperative as long as the forward fuel cell is
empty. Reactivation of the transfer pump
circuits will occur when approximately 18
pounds of fuel enters the forward fuel cell. As
the forward fuel cell is empty at approximately
193.1 pounds of fuel, if the helicopter is shut
down at, or being refueled to between 193.1
and approximately 211.1 pounds of total fuel,
it is possible that up to 18 pounds of fuel may
remain in the forward fuel cell as unusable.
30 APR 2008—Rev. 4———1-77
BHT-407-MD-1
1-25-P-2-B.
MANUFACTURER’S DATA
S/N 53175 AND SUBSEQUENT
Activation of either annunciator circuit is
dependent on the relationship between the
fuel quantity in the forward fuel tank and the
total quantity of the fuel system.
The L/FUEL XFR or R/FUEL XFR transfer light
(Figure 1-16, Sheet 2) will illuminate when
their respective fuel pressure switch senses a
decreasing boost pump pressure of 1.5 ±0.5
P S I , p ro v i d e d t h e c o n d it io n s s h o w n i n
Table 1-10 (Transfer Light Activation Table)
are met:
Table 1-10: Transfer Light Activation Table, S/N 53175 and Subsequent
FWD TANK< 25 POUNDS
FOR 10 CONSECUTIVE
SECONDS
TOTAL FUEL< 250 POUNDS
FOR 10 CONSECUTIVE
SECONDS
L/FUEL XFR OR R/FUEL XFR
LIGHT ILLUMINATED (AT A
DECREASING PRESSURE OF
1.5 ±0.5 PSI)
FALSE
FALSE
YES
FALSE
TRUE
YES
TRUE
FALSE
YES
TRUE
TRUE
NO
With a L/FUEL XFR or R/FUEL XFR pump
pressure of 5 PSI or greater, the respective
left and right fuel pressure switches will open
and cause the L/FUEL XFR or R/FUEL XFR
lights to go off.
When the fuel signal conditioner detects a
forward fuel tank quantity of less than 25
pounds and a total fuel system quantity of
less than 250 pounds, for 10 consecutive
seconds, L/FUEL XFR or R/FUEL XFR transfer
lights will be deactivated to prevent
intermittent light flickering while the
remaining fuel is transferred from the forward
fuel tank to the aft fuel tank.
When forward fuel tank depletion is detected
by the number 3 fuel probe and the low level
detector, the input signals to the fuel signal
conditioner are removed. With the input
signals removed, the fuel signal conditioner
utilizes two 360 second time delays prior to
removing the ground to the right and left
transfer pump relays. This allows the right
and left transfer pumps to continue running
for 360 seconds to ensure all the fuel in the
1-78———Rev. 4—30 APR 2008
forward tank is transferred to the main fuel
tank. The forward fuel cell will be empty when
approximately 193.1 pounds of total fuel is
indicated.
The annunciator and transfer pump circuits
will stay inoperative until the fuel system is
refueled with an appropriate amount of fuel to
reactivate the system. Reactivation of the
tran sfe r pu mp c irc uits w ill oc cu r w h en
approximately 18 pounds of fuel enters the
forward fuel cell. As the forward fuel cell is
empty at approximately 193.1 pounds of fuel,
if the helicopter is shut down at, or being
refueled to between 193.1 and approximately
211.1 pounds of total fuel, it is possible that
up to 18 pounds of fuel may remain in the
forward fuel cell as unusable.
1-25-Q. FUEL FILTER LIGHT
The FUEL FILTER light will illuminate if the
airframe fuel filter is in an impending bypass
condition.
This will occur when a differential pressure of
0.875 ±0.125 PSI is present between the input
MANUFACTURER’S DATA
and output of the fuel filter. The airframe fuel
filter will go into bypass at 3.75 ±0.25 PSI.
BHT-407-MD-1
convert the “TORQUE EVENTS” into
“RETIREMENT INDEX NUMBERS” (RIN) to
track the lives of all required components.
1-25-R. FUEL LOW LIGHT
The FUEL LOW light circuit is designed to
alert the pilot that the main fuel tank quantity
is low. Illuminatio n w ill oc cur w h en
approximately 100 ±10 pounds of usable fuel
remains in the main fuel cell.
A float type low level switch is used to detect
the fuel low condition. The input from the low
level detector is passed through the fuel
signal conditioner which provides a 13 ±3
second time delay to reduce the possibility of
intermittent annunciator flickering due to fuel
sloshing.
In addition, momentary activation (less than 2
seconds) of the FUEL LOW light may be
visible when power is applied to the
helicopter and fuel quantity is greater than the
required low level activation point. This is a
normal occurrence and is a function of fuel
signal conditioner power up logic.
1-26. RETIREMENT INDEX NUMBER
(RIN)
E a c h c o m p o n e n t w it h a r e t i re m e n t li f e
sensitive to "TORQU E EVE NTS" will be
assigned a maximum RETIREMENT INDEX
NUMBER (RIN). This RIN corresponds to the
maximum allowed fatigue damage resulting
from lifts and takeoffs. A new component will
begin with an accumulated RIN of zero that
will be increased as lifts and takeoffs are
performed. The operator will record the
number of lifts and takeoffs and increase the
accumulated RIN accordingly. When the
maximum RIN is reached, the component will
be removed from service. Certain
components may be assigned a life in hours
in addition to the RIN.
Pilots are to record “TORQUE EVENTS” for
each flight. Maintenance personnel will
A "TORQUE EVENT" is defined as a takeoff
(one takeoff plus the subsequent landing =
one RIN) or a lift (internal or external). For
example, if an operator performs six takeoffs
and ten sling loads, this would total 16 torque
events: (6 takeoffs = 6 events, 10 sling loads =
10 events, 6 + 10 = 16 events total).
1-27. TRANSMISSION
The transmission and mast assembly
(Figure 1-18) transfers the engine torque to
the main rotor system with a two stage gear
reduction of 15.29 to 1.0 (6317 to 413 RPM).
The transmission assembly is made up of a
top support case and lower case which
contains an input pinion and bevel gear
arrangement, a planetary gear train, and an
accessory gear drive. The components that
are attached to the transmission and mast
assembly are the engine to transmission
d r i v e s h a ft , t r a n s m i s s i o n o i l p u m p ,
transmission oil filter housing, hydraulic
pump, rotor RPM monopole pickup, and two
electric chip detectors.
The transmission assembly is attached to the
roof of the helicopter, forward of the engine
by a pylon installation. The pylon installation
uses two side beams, four elastomeric corner
mounts, and two for/aft restraint springs.
1-27-A. FREEWHEEL ASSEMBLY
The freewheel unit (Figure 1-18) is mounted
on the engine gearbox, and is driven under
power from the engine power takeoff gear
shaft. Engine power is transmitted to the
outer race of the freewheel unit, then through
the engaged sprag clutch which drives the
freewheel inner shaft and couples the engine
to the transmission input drive shaft.
31 JAN 2007—Rev. 2———1-79
BHT-407-MD-1
MANUFACTURER’S DATA
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Filler cap
Corner mount
Mast assembly
Monopole sensor (NR)
Pylon beam assembly
Upper chip detector
Transmission assembly
Engine-to-transmission driveshaft
Disc (standard)
Freewheel assembly
Freewheel chip detector
Rotor brake kit (optional)
Lower chip detector
10
9
3
8
4
1
2
5
6
11
7
12
13
407_MD_01_0018
Figure 1-18. Transmission Assembly
1-80———Rev. 2—31 JAN 2007
BHT-407-MD-1
The tail rotor drive system is driven through a
flex ible co up lin g an d a splin ed a dap ter
mounted on the aft end of the freewheel inner
shaft.
D u r i n g a u t o r o ta t i o n , t h e s p r a g c l u t c h
disengages and the rotational forces of the
main rotor are allowed to drive the
transmission mounted accessories, and tail
rotor drive system.
1-27-B. TRANSMISSION OIL SYSTEM
The transmission oil system (Figure 1-19)
l u b r i c a t e s t h e t r a n s m i s s io n , m a s t , a n d
freewheel assemblies. An oil level sight gauge
is located on the right side of the transmission
lower case and may be viewed through a
cutout in the air induction cowling. A nonvented filler cap is located on the right side of
the top case to fill the transmission oil system.
The transmission accessory gear drives the
oil pump which delivers 6.0 to 6.7 GPM. The
pressure is controlled by a pressure regulator
valve which is set at approximately 52 PSI. The
pump scavenges oil from the lower case sump
through a wire screen and the lower chip
detector. It is then directed to the transmission
mounted oil manifold and the filter element.
To ensure oil flow is not restricted, the filter
incorporates an impending bypass indicator
button on the end of the filter housing which
will extend at 14 ±2 PSI. Ensure appropriate
maintenance actions are carried out following
an impending bypass indication in accordance
with the BHT-407-MM. Additionally, a filter
bypass valve will open if the differential
pressure reaches 17 PSI and if the differential
pressure reaches 29.6 to 38.6 PSI, a high
pressure relief valve will open and the oil will
totally bypass the filter.
The transmission mounted oil manifold also
incorporates a thermostatic valve which
MANUFACTURER’S DATA
controls the flow of oil to the oil cooler. At oil
temperatures below 150°F (66°C), the oil
cooler is bypassed and the oil is directed back
to the transmission. As the oil temperature
increases above 150°F (66°C) oil is gradually
directed to the oil cooler until at 180°F (82°C),
all oil is directed through the cooler.
A temperature bulb for the oil temperature
indicator and a thermoswitch for the XMSN
O I L T E M P l ig h t a r e a l s o l o c a t e d o n t h e
transmission mounted oil manifold.
After the oil exits the cooler (or bypasses the
cooler through the thermostatic bypass valve),
it is then directed to the transmission to
lubricate the various gears and bearings. The
oil is then directed to a deck mounted oil
manifold located below the transmission input
driveshaft. A pressure transducer and oil
pressure switch are mounted on the
transmission deck oil manifold. The
transducer provides signals to the oil pressure
gauge and the pressure switch controls the
XMSN OIL PRESS light.
After leaving the deck mounted oil manifold,
the oil flows through the forward fire wall and
a tee fitting equipped with two restrictors. The
restrictor fittings reduce the flow and direct
the oil to the freewheel forward duplex bearing
and aft housing bearing. The oil that lubricates
the bearing in the aft housing moves forward
through the hollow engine output driveshaft to
the freewheel sprag clutch and bearing. The
oil is then collected in the forward freewheel
housing and returned to the main
transmission lower case at atmospheric
pressure.
1-27-C. TRANSMISSION INDICATORS
Tr a n s m i s s i o n i n d i c a t o r s i n c l u d e a n o i l
temperature and pressure gauge, XMSN OIL
TEMP light, XMSN OIL PRESS light, and XMSN
CHIP light.
31 JAN 2007
Rev. 2
1-81
BHT-407-MD-1
MANUFACTURER’S DATA
PRESSURE
REGULATOR
(REGULATES OIL
PRESSURE FROM
50 TO 55 PSI)
OIL FILTER DIFFERENTIAL PRESSURE
BY-PASS VALVE AND BY-PASS
INDICATOR BUTTON (ACTUATION
PRESSURE = 14 ±2 PSID AND VALVE
CRACKING PRESSURE = 17 PSID MIN.)
THERMOSTATIC
BY-PASS VALVE
(CRACKS AT 190°F
MIN. AT 40 PSID
MIN.) (STARTS
CLOSING AT 150°F
AND IS CLOSED
AT 178°F ± 2°F
SUN GEAR OIL FLOW
(3 JETS)
PLANETARY OIL FLOW
(3 JETS)
SUN GEAR SPLINE
OIL PRESSURE SWITCH
(ACTUATES AT 32 +2 PSI - LOW PRESSURE
GOING DOWN PRESSURE ACTUATES THE LIGHT
AT 38 PSI - HI PRESSURE DEACTUATES THE
LIGHT PRESSURE INCREASING)
MAIN TRANSMISSION
AND MAST ASSEMBLY
OIL JET 4 SPIRAL
BEVEL INTO MESH
(NOT SHOWN)
OIL JET 2
MAST BEARING
(ANNULUS-SLOT)
PLANETARY OIL FLOW
OIL PRESSURE TRANSMITTER
TEST PORT
OIL FILTER CAP
MANIFOLD
OIL JET 3 SPIRAL
BEVEL INTO MESH
(NOT SHOWN)
OIL SUMP
RESTRICTOR FITTING ASSEMBLY
0.027 TO 0.032 IN. RESTRICTOR HOLE
OIL
COOLER
OIL FILTER
2HIGH PRESSURE
RELIEF VALVE
(CRACKS AT 29.6
TO 38.6 PSI)
RESTRICTOR FITTING ASSY
0.035 TO 0.040 IN. RESTRICTOR HOLE
OIL FILTER
TEE FITTING
TEMP BULB
OIL
PUMP
THERMOSWITCH
ACTUATES AT 230°F
OIL
DRAIN
OIL SUMP
OIL LEVEL
SIGHT GAUGE
DUPLEX BEARING
OIL FLOW
(ANNULUS - SLOT)
CHIP DETECTOR AND
OIL INLET SCREEN
OIL JET 1
ROLLER BEARING-SPIRAL
BEVEL PINION SUPPLIES OIL
FLOW TO OIL JETS 3 AND 4
LEGEND
FREEWHEEL INSTALLATION
Supply oil at pump pressure, 80 to 150 PSI.
NOTE
Supply oil at regulated delivery pressure, 50 to 55 PSI.
OIL PUMP
DESIGN FLOW RATE = 6.0 TO 6.7 GPM
MAXIMUM PRESSURE = 150 PSI
RATED PRESSURE = 80 PSI
Return oil at atmospheric pressure.
Scavenge oil
407_MD_01_0019
Figure 1-19. Transmission Oil System
1-82
TRIPLEX BEARING
OIL FLOW
(ANNULUS - SLOT)
Rev. 2
31 JAN 2007
MANUFACTURER’S DATA
1-27-D. TRANSMISSION OIL
TEMPERATURE AND PRESSURE
GAUGE
The transmission oil temperature and
pressure gauge is a dual instrument that
simultaneously displays oil temperature in
degrees Celsius on the right side display and
oil pressure in PSI on the left side display.
Each side of the indicator is powered by its
own circuit breaker.
The transmission temperature input signal is
provided by a thermobulb installed on the
transmission oil filter manifold. The
transmission oil pressure is provided by a
transducer mounted on the transmission
deck oil manifold.
1-27-D-1.
TRANSMISSION OIL
TEMPERATURE GAUGE
The XMSN OIL TEMP light will be illuminated
when the transmission oil temperature switch
detects a temperature of 110 ±5.6°C (230
±10°F). The oil temperature switch is mounted
o n th e t r a n s m i s s io n o il f i l t e r m a n i f o l d
housing.
1-27-D-2.
TRANSMISSION OIL PRESSURE
LIGHT
The XMSN OIL PRESS light will be illuminated
when the transmission oil pressure switch
detects a decreasing oil pressure of 30 ±2 PSI.
Similarly, as the transmission oil pressure
builds, the oil pressure switch will extinguish
the XMSN OIL PRESS annunciator prior to the
pressure passing through 38 PSI.
BHT-407-MD-1
Two magnetic chip detectors, one on the
upper case and one on the lower case, are
mounted on the main transmission. The lower
case chip detector incorporates a self-sealing
valve which prevents the loss of oil from the
gearbox when the chip detector is removed.
The third chip detector is mounted on the
forward freewheel housing and also
incorporates a self-sealing valve which
prevents the loss of oil from the freewheel
unit when the chip detector is removed.
1-28. ROTOR SYSTEM
1-28-A. MAIN ROTOR HUB AND BLADES
The rotor assembly (Figure 1-20) is a four
bladed soft-in-plane design with a 35 foot
diameter rotor.
The main rotor hub contains a glass/epoxy
compo site yo ke tha t acts as a flap ping
flexure. Elastomeric bearings and dampers
which require no lubrication are utilized. The
hub also incorporates the use of lead-lag,
coning/flapping and droop stops.
The main rotor blades are a composite design
utilizing a glass/epoxy spar, glass/epoxy
skins, and a nomex core afterbody. The
blades incorporate a nickel plated stainless
steel leading edge erosion strip and are
coated with conductive paint for lightning
protection. The blades are also individually
interchangeable.
1-28-B. TAIL ROTOR HUB AND BLADES
1-27-E. TRANSMISSION CHIP
ANNUNCIATOR
The tail rotor (Figure 1-21) is a two-bladed
teetering rotor with a 5.42 foot diameter. It is
mounted on the left side of the tailboom and
rotates clockwise when looking inboard from
the left side of the helicopter.
The XMSN CHIP light will illuminate if
magn etic par tic les in th e oil sy stem
accumulate on any one of the three quick
disconnect magnetic chip detectors.
Teflo n li ne d p itch ch an g e be ar ing s a re
installed in a steel yoke assembly which
utilizes an elastomeric flapping bearing and
flapping stops.
31 JAN 2007—Rev. 2———1-83
BHT-407-MD-1
MANUFACTURER’S DATA
1
2
4
14
16
15
17
3
6
13
5
8
12
7
1 11
9
10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
Cover
Frahm damper
Safety lock
Main rotor blade grip
Yoke
Damper
Elastomeric lead-lag bearing
Pitch link
Mast
Lower cone
Center cone set (if applicable)
Pitch horn
Elastomeric shear bearing
Coning/flapping stop
Upper plate
Mast nut
Upper cone
NOTE
1
Center cone set applicable to S/N 53000 through 53631 Pre TB 407-05-66. S/N 53632 and subsequent
have an integral mast center cone.
407_MD_01_0020
Figure 1-20. Main Rotor Assembly
1-84———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
10
BHT-407-MD-1
12
11
1
9
2
7
SEE
VIEW A
8
6
3
5
4
1
1
0.055 IN. (1.39 mm)
MINIMUM YIELD
INDICATOR DIMENSION
DEFORMED
YIELD INDICATOR
VIEW A
STATIC STOP
CROSS SECTION
1.
2.
3.
4.
5.
6.
Tail rotor blade
Tail rotor pitch change link
Crosshead
Counterweight assembly
Tail rotor hub
Pitch horn
7.
8.
9.
10.
11.
12.
Static stop
Tail rotor gearbox
Chip detector
Pitch change mechanism
Breather line
Tail rotor gearbox filler cap
NOTE
1
S/N 53351 and subsequent, and helicopters modified per ASB 407-99-27 or TB 407-99-17.
407_MD_01_0021
Figure 1-21. Tail Rotor Assembly
31 JAN 2007—Rev. 2———1-85
BHT-407-MD-1
The blades are a composite design utilizing a
glass/epoxy spar, glass/epoxy skins, and a
nomex core. The blades incorporate nickel
plated stainless steel leading edge abrasion
strip and are coated with conductive paint for
lightning protection.
The tail rotor yoke static stop has been
designed with yield indicators. The yield
indicators provide the ability to visually
determine if the tail rotor yoke has been
stressed beyond designed limits. This will be
evident by deformation of either of the static
stop yield indicators due to excessive contact
with the yoke (Figure 1-21). If deformation of
either yield indicator is evident, contact
maintenance personnel prior to further flight.
1-28-C. TAIL ROTOR GEARBOX
The tail rotor gearbox (Figure 1-21), located
on the aft end of the tailboom, drives the tail
rotor. It contains two spiral bevel gears
positioned at 90° angles to the other. The tail
rotor gear box has a gear reduction of 2.53 to
1.0 which reduces the driveshaft input speed
of 6317 RPM to an output shaft speed of 2500
RPM.
The gearbox has a self contained oil
lubrication system, non vented filler cap, and
a magnetic chip detector.
1-28-D. TAIL ROTOR CHIP LIGHT
The T/R CHIP light will illuminate if magnetic
pa r t i c l e s i n t h e o i l a c c u m u l a t e o n t h e
magnetic chip detector.
MANUFACTURER’S DATA
The NP side of the instrument is powered by
th e 28 VDC bus thro ugh its own circuit
breaker.
An NP Speed Pickup mounted on the engine
provides two separate signals to the FADEC/
ECU. One signal is the primary driver of the
N P indicator and is also a secondary signal
for the FADEC/ECU. The second signal is the
primary input to the FADEC/ECU. The primary
signal to the indicator will be passed through
the FADEC/ECU even when the FADEC/ECU is
not powered.
The NR side of the instrument is powered by
th e 28 VDC bus thro ugh its own circuit
breaker.
An NR speed pickup mounted on the
transmission lower case provides three
separate identical signal outputs of rotor
RPM. One signal is sent to the NR circuit of
the dual tachometer gauge. One signal is sent
to the FADEC/ECU and the other is sent to the
Rotor RPM sensor switch.
1-29-B. RPM LIGHT
HORN
AND
WARNING
The RPM light and warning horn circuit is
designed to activate when the main rotor RPM
(NR) is less than 95%. The RPM light will also
be illuminated at a N R speed of 107% and
higher.
1-29-B-1.
NR LESS THAN 95%
1-29. ROTOR SYSTEM INDICATORS
If the rotor RPM sensor detects a main rotor
RPM (NR) of less than 95%, it will illuminate
the RPM light and (continuous sounding) low
rotor RPM horn. The low rotor RPM horn can
be muted by pressing the warning horn mute
switch.
1-29-A. DUAL TACH GAUGE
1-29-B-2.
The Dual Tachometer has two pointers that
simultaneously display Rotor RPM (NR) on the
outer scale and the engine N P on the inner
scale. All scales are in percent RPM.
If the rotor RPM sensor detects a main rotor
RPM NR of 107% or greater, it will illuminate
the RPM light. The low rotor RPM horn will not
be activated.
The chip detector incorporates a self-sealing
valve which prevents the loss of oil from the
gearbox when the chip detector is removed.
1-86———Rev. 2—31 JAN 2007
NR 107% OR GREATER
BHT-407-MD-1
1-30. FLIGHT CONTROL SYSTEM
1-30-A. ROTOR CONTROLS
Main rotor and tail rotor flight control systems
(Figure 1-22), consisting of cyclic, collective
and anti-torque controls are used to regulate
the helicopter attitude, altitude and direction
of flight. The flight controls are hydraulically
boosted to reduce pilot effort and to
counteract control feedback forces.
1-30-B. MAIN ROTOR
Main rotor cyclic and collective flight controls
regulate pitch and roll attitude and thrust.
Control inputs from the cyclic and collective
control sticks (Figure 1-22) in the cockpit are
transmitted by push-pull tubes to hydraulic
servo actuators mounted on the top deck. The
actuators operate the cyclic and collective
l e v e r s , w h i c h r a i s e , l o w e r, a n d t i l t t h e
swashplate. The swashplate converts fixed
control inputs to the rotating controls and
allows cyclic and collective pitch inputs to the
main rotor.
In the case of loss of hydraulic pressure to the
servo actuators, springs are installed in
parallel to the cyclic and collective push-pull
tubes on the cabin roof to assist the pilot with
the increased control feedback forces.
1-30-B-1.
CYCLIC
The cyclic control stick (Figure 1-22) is
mounted under the pilots crew seat and
protrudes from the forward bulkhead of the
crew seat. The fore and aft cyclic input is
connected through push-pull tubes to the
cyclic hydraulic servo actuators. In addition,
all cyclic fore and aft movement is fed through
a cam assembly that automatically adds an
a m o u n t o f l a t e r a l c y c l i c in p u t t h a t i s a
p e r c e n ta g e o f t h e f o r e a n d a f t c y c l i c
movement. A spring canister is provided in
line with th e cam inp ut to pe rm it cyclic
MANUFACTURER’S DATA
movement in the event that the cam assembly
becomes jammed.
The lateral cyclic input is connected through
push pull tubes to the cyclic hydraulic servo
actuator. The hydraulic servo actuators
operate bellcranks and push pull tubes that tilt
t h e s w a s h p l a t e n o n - r o ta t i n g r i n g . T h e
swashplate rotating ring tilts likewise and
actuates the pitch links which control the
plane of rotation of the main rotor.
The cyclic control stick grip contains a twoposition intercommunication/radio transmit
switch and a cargo hook release switch. An
adjustable friction control knob, located at the
base of the cyclic stick where it protrudes
through the forward crew seat bulkhead,
allows the pilot to set the desired amount of
control stiffness for flight or to lock the cyclic
control stick during ground operation or
shutdown.
A cyclic stick position switch (cyclic centering
s w i t c h ) i s a t ta c h e d t o t h e c y c l i c s t i c k
bellcrank. When the helicopter is on the
ground, this switch will cause the CYCLIC
CENTERING annunciator in the caution/
warning panel to illuminate when the stick is
not centered.
1-30-B-2.
COLLECTIVE
The collective control stick (Figure 1-22) is
mounted between the pilot and copilot crew
seats. The collective control stick controls the
collective hydraulic servo actuator through
push-pull tubes. This operates the collective
lever mounted on the top of the transmission.
The collective lever raises and lowers the
swashplate ball-sleeve assembly and the
cyclic levers to induce collective pitch to the
main rotor blades without affecting the cyclic
path. A spring is installed under the copilot's
crew seat to balance the required force to
raise and lower the collective with the
hydraulic boost system operating.
31 JAN 2007
Rev. 2
1-87
BHT-407-MD-1
MANUFACTURER’S DATA
LDG
LTS
BOTH
MAIN ROTOR
PITCH LINK
START
F
W
D
FLOAT
ARM
OFF
FLOAT
INFLATE
SWASHPLATE
ASSEMBLY
DISENG
ROTATING RING
NON ROTATING RING
11
10
DETAIL B
CYCLIC GRIP
12
11
9
DETAIL A
PILOT COLLECTIVE STICK
SEE
DETAIL
B
SEE
DETAIL A
2
1
8
7
6
5
13
ALTERNATE
PEDALS
15
3
16
4
1
14
17
20
Pilot tail rotor control pedal assembly
Pilot cyclic stick
Cyclic friction knob
Cyclic lateral balance spring
Cyclic centering switch
Cam override spring
Pilot collective stick
Collective friction knob
Collective servo actuator
Cyclic servo actuators
Balance springs
Tail rotor servo actuator
Collective balance spring
Copilot collective stick
Collective pitch transducer (FADEC SYS)
Cyclic cam
Cyclic longitudinal balance spring
Copilot cyclic stick
Copilot tail rotor control pedal assembly
Control pedal spring
NOTE
18
TYPICAL
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
1
Alternate pedals installed in production at S/N 53730 and subsequent.
19
407_MD_01_0022
Figure 1-22. Flight Controls
1-88
Rev. 2
31 JAN 2007
MANUFACTURER’S DATA
A collective friction knob is located near the
base of the collective stick between the pilot
and copilot seats. A throttle twist grip for the
engine is mounted on the collective stick. A
me chan ical idle re le ase pus h bu tto n is
located in front of the twist grip throttle. A
switchbox located on the forward end of the
collective stick provides a base for the engine
start switch and landing light switch.
1-30-C. TAIL ROTOR
The tail rotor, or anti-torque, flight controls
(Figure 1-22) provide pitch adjustment of the
tail rotor blades for yaw control. A set of
pedals on the cockpit floor, forward of the
pilot seat, are connected to a directional
control hydraulic servo actuator, located in
the aft fuselage near the tailboom. Push-pull
tubes connect the actuator to the fixed pitch
change mechanism of the tail rotor gearbox.
The tail rotor fixed mechanism is connected
to the rotating controls through a rotating
push-pull tube. The push-pull tube attaches to
a sliding crosshead that moves in and out on
splines on the tail rotor mast to provide pitch
control. Rotating counterweights minimize the
control forces required.
T h e ta i l ro to r c o n t r o l p e d a ls c o n ta in a
bellcrank pedal adjuster, which provides for
manual adjustment of pedal position
according to the pilot’s needs. Alternate
pedals may also be installed which allow the
pilot to manually adjust the position of the
pedal foot rests.
For helicopters with dual controls, the copilot
fully fu nc tion al ta il ro tor co ntr ol ped al
assembly is installed on the floor in front of
the copilot seat to provide a means for the
copilot to control the tail rotor assembly. The
control pedals are linked to the pilot pedals by
means of control tubes and a bellcrank. The
copilot pedals can also be positioned, as
desired, by means of the pedal adjuster.
Alternate pedals may also be installed which
allow the copilot to manually adjust the
position of the pedal foot rests.
BHT-407-MD-1
1-30-D. AIRSPEED ACTUATED PEDAL
STOP SYSTEM
The Airspeed Actuated Pedal Stop system is
comprised of a Pedal Restrictor Control Unit
( P R C U ) , a n a c t u a t i n g r o ta r y s o l e n o i d ,
positioning sensing microswitch, and a
press-to-test PEDAL STOP (PTT) switch.
A PEDAL STOP segment (Figure 1-4) on the
C a u t i o n / Wa r n i n g p a n e l i n d i c a t e s a n y
malfunction of the system. System power is
provided through a 5-amp circuit breaker
located in overhead console (Figure 1-5).
The helicopter pitot and static installation
interfaces with the PRCU. The PRCU
calculates airspeed from the pitot and static
inputs and when greater than 55 ±5 KIAS,
drives the solenoid to extend the pedal stop
restrictor into the left pedals range of travel.
The PRCU will trigger the solenoid to retract
the pedal stop when calculated airspeed falls
below 50 ±5 KIAS.
Upon full extension of the pedal stop, the
position sensing microswitch is activated and
the PRCU illuminates the ENGAGED message
on the PED AL STO P (P TT) s witch. This
message is extinguished when the pedal stop
is retracted.
The Airspeed Actuated Pedal Stop System
diagram is shown in Figure 1-23.
1-31. HYDRAULIC SYSTEM
The hydraulic system (Figure 1-24) provides
boost power for the cyclic, collective and
an ti-to rqu e flig ht c on tr ols. The s yste m
includes a pump, reservoir, pressure and
return filter assemblies, pressure and return
m a n i fo ld , p re s s u re m o n it or i ng s e n s o r,
solenoid valve, pressure relief valve, flight
control servo actuators, and interconnecting
tubing and fittings.
The hydraulic pump is mounted on and driven
by the transmission. The pump is a variable
delivery
pressure
compensated,
31 JAN 2007—Rev. 2———1-89
BHT-407-MD-1
self-lubricated type designed to operate
continuously and provide a rated discharge
pressure of 1000 -25/+50 PSI. Hydraulic fluid
is supplied to the pump from a vented gravity
feed reservoir mounted forward of the
transmission. Fluid passes through the pump,
a pressure filter, and a solenoid valve. The
solenoid valve is controlled by a HYD SYS
switch located on the overhead console.
When HYD SYS switch is ON, the solenoid
valve is de-energized to the open position and
fluid is routed to cyclic, collective, and
anti-torque actuators. From the actuators,
fluid passes through a return filter to the
reservoir. Both the pressure and return filter
assemblies have filter indicators. These
indicators are activated when the differential
pressure across the filter is 70 ±10 PSI. The
pressure filter does not contain a bypass
valve. The pressure filter will clog completely
in order to prevent contaminated hydraulic
fluid from being pumped through the system.
The return filter contains a bypass valve that
will allow fluid to bypass the filter if it senses
a differential pressure of 100 ±25 PSI across
the filter. When the differential pressure
decreases to 60 PSI, the bypass valve will
close. This allows fluid returning from the
servo actuators to return to the reservoir even
if the return filter is completely clogged.
A relief valve is incorporated in the system,
between the pressure filter and the solenoid
valve. The relief valve is normally closed.
When the pressure reaches 1225 ±150 PSI, the
relief valve will open to protect the system
from damage. The relief valve will reset when
pressure drops to approximately 1075 PSI.
1-31-A. HYDRAULIC INDICATORS
Hydraulic system indicators include a
HYDRAULIC SYSTEM caution light on the
caution/warning panel and filter bypass
indicators located on both the pressure and
return filter assemblies.
1-90———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
1-31-B. HYDRAULIC FILTER
INDICATORS
Each filter assembly contains a filter indicator
that indicates an impending clogged filter.
T h e i n d i c a t o r c o n s i s ts o f a r e d b u t t o n
mounted on the filter assembly housing.
When the differential pressure across the
filter is 70 ±10 PSI, the red button will rise. To
prevent inaccurate indications of bypass, the
indicator will not work when the hydraulic
fluid temperature is less than 35°F (2°C). If
hydraulic fluid temperature is more than 35°F
(2°C), the indicator gives the correct
indication of clogging, even if the ambient
temperature is below 35°F.
If a filter button is extended, indicating an
impending clogged filter, maintenance action
should be performed prior to next flight. The
filter indicator is reset by pushing the button in.
1-31-C. HYDRAULIC SYSTEM LIGHT
T h e H Y D R A U L I C S Y S T E M l ig h t c i rc u i t is
designed to illuminate when the hydraulic
system fluid pressure is below the minimum
limit.
The hydraulic system light will be illuminated
when the hydraulic pressure switch detects a
decreasing pressure of 650 -0/+100 PSI, and
will extinguish on an increasing pressure at
750 +0/-100 PSI. The hydraulic pressure
switch is mounted on the hydraulic manifold,
forward of the hydraulic actuator support.
1-32. ELECTRICAL SYSTEM
The helicopter is equipped with a 28 VDC
electrical system (Figure 1-25). Power for this
system is obtained from a nickel-cadmium 24
volt, 17 amp/hour battery or optional 24 volt,
28 amp/hour battery and a 30 volt, 200-amp
starter-generator. The starter-generator has
been derated to 180 amps to ensure adequate
cooling under all operating conditions up to
18,000 feet H p . Refer to the BHT-407-FM
limitations for operations above 18,000 feet
Hp .
BHT-407-MD-1
MANUFACTURER’S DATA
PEDAL STOP (PTT)
SWITCH
PEDAL STOP
PTT
LTG POWER
28/15 VDC
ENGAGED
LTG POWER
28/15 VDC
6540CB1
PRCU
28 VDC PWR
TEST INPUT
CAUTION
WARNING PANEL
POSITION ANN GND
PEDAL STOP
PTT
FAULT ANN
SOLENOID DRIVER
POSITION IND
POSITION IND RTN
POWER GND
FAULT ANN GND
PITOT
INPUT
STATIC
INPUT
407_MD_01_0023
Figure 1-23. Airspeed Actuated Pedal Stop System
31 JAN 2007
Rev. 2
1-91
BHT-407-MD-1
MANUFACTURER’S DATA
OPERATING PRESSURE 1000 PSI -25/+50
HYDRAULIC FLUID MIL-H-5606
SEE DETAIL B
SEE DETAIL B
16
1
2
3
17
SEE DETAIL B
4
6
5
15
15
7
8
CYCLIC
9
HYDRAULIC SYSTEM
14
10
5
SERVO
PRESSURE
ACTUATOR
DE-ENERGIZEDSYSTEM ON
TAIL ROTOR
DIFFERENTIAL
RELIEF VALVE
TEST PORT
8
PRESSURE
RETURN
RETURN
RETURN
PRESSURE
CAUTION LIGHT
PRESSURE
RETURN
SUCTION
11
COLLECTIVE
SEE DETAIL A
13
12
CYCLIC
18
SERVO
ACTUATOR
SEQUENCE
VALVE
ENERGIZEDSYSTEM OFF
SOLENOID VALVE SCHEMATIC
CHECK
VALVES
INPUT FROM
FLIGHT CONTROLS
DETAIL A
17.
18.
Tail rotor servo actuator
Hydraulic pressure switch
CYLINDER
OUTPUT
SERVO ACTUATOR - TYPICAL
SEE DETAIL B
407_MD_01_0001
Figure 1-24. Hydraulic System
1-92
Rev. 2
31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
EXTERNAL
POWER RECEPTACLE
BATTERY
-
BATTERY
RELAY
+
EXTERNAL
POWER RELAY
+
+
OFF
-
BATTERY
SWITCH
BAT (ON)
AMMETER
AVIONICS SWITCH
28 VDC BUS
STARTER
RELAY
GENERATOR
RELAY
BUS BAR
BUS BAR
AVIONICS BUS
-
+
SHUNT
STARTER
GENERATOR
C
D
GEN FIELD GEN FIELD
E
10
GENERATOR
SWITCH
A
B
FADEC/START
RELAY
15
START
SWITCH
RESET
START
GEN RESET
OFF
5
DC CONTROL UNIT/
VOLTAGE REGULATOR
J
L
M
D
G
K
GEN
FADEC
MODE SWITCH
DISENG
AUTO
FADEC/ECU
START (-)
AUTO
MANUAL
H
ENG START
MANUAL
5
TYPICAL
407_MD_01_0024
Figure 1-25. DC Electrical System
31 JAN 2007—Rev. 2———1-93
BHT-407-MD-1
Major components of DC power system
include battery, starter-generator, DC control
unit, voltage regulator, relays, 28 VDC bus,
and circuit breakers. All circuits in electrical
system are single wire with fuselage common
ground return. Negative terminals of
starter-generator and battery are grounded to
helicopter structure. Controls for electrical
system are located on overhead console and
instrument panel.
Generator is provided with over voltage,
under voltage and reverse current protection.
If an over voltage (32 ±0.5 VDC), under voltage
(18 ±1.8 VDC), or reverse current (0.08 to 0.150
VDC for 350 milliseconds) is detected, the DC
control unit/voltage regulator will disconnect
the generator from the system. Failure of the
generator can be determined by GEN FAIL
light, a zero ammeter reading, and battery
voltage displayed on the voltmeter.
In the event power from the generator is lost,
emergency power available from the battery
can be maximized by pulling the circuit
breakers on all non-essential systems. In the
event of a total electrical failure (hard short on
bus), if the battery switch is immediately
placed to the OFF position, the 17 amp/hour
battery, assuming 80% charged, can supply
fuel boost and transfer pumps for a period of
approximately 1.7 hours. The optional 28 amp/
hour battery, assuming 80% charged, can
supply fuel boost and transfer pumps for a
period of approximately 2.8 hours.
1-32-A. EXTERNAL POWER
Ex te rna l po w er ma y b e su pp lie d to th e
helicopter by means of a receptacle located
on the lower front section of helicopter. 28
VDC Ground Power Unit (GPU) shall be 500
amps or less to reduce risk of starter damage
from overheating.
If external power was used to power the start
and the battery switch was left in the OFF
position, it is important to position the battery
switch to ON prior to removing the external
power source (refer to BHT-407-FM-1, Normal
1-94———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
Procedures). If all sources of electrical power
are removed from the ECU with the engine at
idle in AUTO mode, the start solenoid valve in
the HMU will open, causing the engine to
decelerate and possibly flame out. If the
Battery switch is inadvertently left OFF and
the external power source is removed, do not
attempt to reapply power when a decrease in
NG speed is noted. Throttle should be
positioned to cutoff. Reapplication of
electrical power could cause an over
temperature condition due to the reduced NG
speed and reintroduction of fuel by the
FADEC system.
1-32-B. BATTERY SWITCH
T h e b a t t e r y s w i t c h i s i n s ta l l e d o n t h e
overhead console and controls the battery
relay which connects the battery to the DC
bus. The switch has two positions OFF and
BATT.
In the event BATTERY HOT light comes ON,
BATT switch shall be turned OFF. If BATT
switch is turned OFF and BATTERY RLY light
illuminates, this indicates battery relay
contacts have not opened. If battery relay
c o n ta c ts h a v e n o t o p en e d , b a tt e ry w ill
continue to receive a charge from the
generator. To prevent this, pilot should turn
GEN switch to OFF. Battery will continue to
run all electrical systems through the closed
battery relay. To reduce load on overheated
battery, pilot should pull circuit breakers on
all non-essential systems. Left FUEL XFR/
BOOST circuit breaker switch should be left
ON to insure fuel transfer from the forward
fuel tank to the main fuel tank continues.
The BATT switch also configures the DC
power feed to the left fuel boost pump and the
left fuel transfer pump between the DC bus
(battery switch positioned to BATT) and the
helicopter battery (BATT switch positioned to
OFF).
In the event the battery switch is positioned to
OFF during helicopter operations, an
alternate circuit (Figure 1-17) is provided to
MANUFACTURER’S DATA
BHT-407-MD-1
allow operation of the left fuel transfer and left
fuel boost pumps. During this condition, with
the FUEL VALVE switch positioned to ON,
battery voltage is supplied through the fuel
boost/xfr backup circuit breaker, the fuel
valve switch, the battery switch, and the left
fuel xfr/ boost circuit breaker switch to the left
fuel transfer and left fuel boost pumps.
charging is supplied by the GPU and must be
monitored. Charging will be completed when
GPU output indicates approximately 8 amps.
In the event BATTERY HOT warning light
illuminates, battery shall be turned OFF and
GPU power removed.
1-32-C. BATTERY CHARGING (17 AND 28
AMP/HOUR BATTERIES)
The generator switch is installed on the
overhead console and controls generator
output by opening and closing the generator
field circuit. The switch is a double pole,
double throw, spring loaded design with only
momentary contact in the RESET position.
The switch has 3 positions, GEN, OFF, and
RESET.
As a maintenance function, battery charging
may be accomplished with battery installed in
helicopter, due to minor depletion, using a
GPU. The GPU must incorporate a good
quality constant voltage regulator, a variable
voltage selector and an amperage indicator.
NOTE
It is recommended that the charging
procedure not exceed 45 minutes in
duration. Frequent use of this
procedure may cause loss of
electrolyte due to gassing through
electrolysis. R educ ed levels of
electrolyte can cause cell
imbalances and lower overall battery
capacity. This procedure is not
intended to take the place of
scheduled battery maintenance
procedures.
CAUTION
BATTERY HOT CAUTION LIGHT
MUST
BE
CONTINUOUSLY
MONITORED
DURING
THIS
PROCEDURE. IF BATTERY HOT
LIGHT ILLUMINATES, BATTERY
SWITCH SHALL BE SET TO OFF AND
GPU POWER REMOVED TO REDUCE
POSSIBILITY OF THERMAL
RUNAWAY.
Battery switch may be set to ON after GPU
power is applied and voltage adjusted to 28.5
VDC (do not exceed 28.5 volts). Battery
1-32-D. GENERATOR SWITCH
With the generator switch positioned to GEN,
its function is to complete the generator field
circuit between the starter generator and the
generator control unit/voltage regulator.
Under normal operating conditions, this will
allow the generator control unit/voltage
regulator to monitor and control the output
voltage of the starter generator and in turn
connect the output of the generator to the 28
VDC bus through the generator relay.
Positioning the generator switch to OFF
opens the generato r field circuit which
removes control of the generator control unit/
voltage regulator from the generator and
generator relay. The generator relay will open,
removing the generator from the 28 DC bus.
In the event an over voltage condition is
detected by the generator control unit/voltage
regulator, an internal regulator trip relay
circuit will be activated. Positioning the
generator switch to RESET provides bus
power, from the battery, to the generator
control unit/voltage regulator which will reset
the internal trip relay circuit. If the
malfunction condition persists following the
RESET, further attempts to reset should not
be made.
Additionally, following a start using a GPU,
ensure battery switch is positioned to ON and
31 JAN 2007—Rev. 2———1-95
BHT-407-MD-1
GPU is disconnected prior to positioning the
generator switch to GEN. Positioning the
generator switch to GEN with the GPU
connected may cause a reverse current
situation and trip the generator off line.
1-32-E. START SWITCH
The start switch is located on the collective
switch box. It contains two sets of spring
loaded contacts which provide momentary
con tact in either th e S TA RT or D IS EN G
positions. When the switch is positioned to
start or disengage, only one set of contacts
move (Figure 1-25). For a description of a
START IN AUTO MODE or START IN MANUAL
MODE refer to paragraph 1-12-D or paragraph
1-12-F.
1-32-F. ELECTRICAL SYSTEM
INDICATORS
Electrical system indicators include a DC
ammeter, voltmeter, BATTERY HOT light,
BATTERY RELAY light, START light, and GEN
FAIL light.
MANUFACTURER’S DATA
Amperes 301 to 400 amps — Maintain full
scale for 5 seconds. After 5 seconds,
drop to zero.
If ammeter indication drops to zero,
p o s i t i o n g e n e r a t o r s w i t c h t o O F F.
Following a brief time, generator switch
can be positioned to GEN (ON). Refer to
the BHT-407-FM-1 for generator load
limitations.
1-32-G-2.
AMMETER — 400 AMPS MAX
SCALE
Ammeter indications will be continuous
regardless of load. Indicator will not drop to
zero if limits are exceeded. Refer to the
BHT-407-FM-1 for generator load limitations.
To ensure limits are not exceeded, pilot can
switch generator to OFF prior to generator
exceedance being reached. Following a brief
time, generator switch can be positioned to
GEN (ON).
1-32-H. VOLTMETER
1-32-G. FUEL PRESSURE/DC AMMETER
The fuel pressure/ammeter gauge is a dual
display instrument. The ammeter indicates
the load in amperes that is being supplied to
the 28 VDC bus by the engine driven
generator.
The ammeter side of the indicator is powered
by its own circuit breaker.
1-32-G-1.
AMMETER — 200 AMPS MAX
SCALE
Fo r a m m et er re a din g s a bo ve 20 0 a m ps
(maximum scale) the indicator will indicate as
follows:
Amperes 201 to 300 amps — Maintain full
scale for 2 minutes. After 2 minutes, drop
to zero.
1-96———Rev. 2—31 JAN 2007
The voltmeter is included in a multifunction
indicator mounted in the upper left area of the
instrument panel. The indicator also displays
Outside Air Temperature and Clock functions.
A button located on the center top of the
instrument changes the top display between
v o l ts ( e . g . , 2 8 E ) a n d O AT ( C e l s i u s a n d
Fahrenheit). When power is applied to the
i n s t r u m e n t t h e d i s p l a y d e f a u l ts t o t h e
voltmeter reading.
The voltmeter display receives its power from
the 28 VDC bus through the OAT/V INSTR
circuit breaker. The 28 VDC bus through this
circuit breaker is also the source of the
voltage value read and displayed by the
voltmeter. When all electrical power is turned
off, the voltmeter display disappears.
MANUFACTURER’S DATA
1-32-I.
BATTERY HOT ANNUNCIATOR
The BATTERY HOT light is utilized with either
the basic ship 17 amp/hour or optional kit 28
amp/hour battery installed. The 17 amp/hour
battery incorporates two thermal switches
while the 28 amp/hour battery incorporates
three thermal switches. While both of the 17
amp/hour battery thermal switches are used
in the BATTERY HOT light circuit, only two of
the three thermal switches available on the 28
amp/hour battery are used.
The 17 amp/hour battery thermal switches will
close at a temperature of 145 ±5°F (62.7
±2.8°C). The 28 amp/hour battery switches will
close at a temperature of 160 ±5°F (71.1
±2.8°C).
When any one of the thermal switches closes,
the BATTERY HOT light will be ON.
1-32-J. BATTERY RELAY
ANNUNCIATOR
The BATTERY RELAY light will be ON if the
battery relay has remained in the closed
(energized) position after the battery switch
has been set to OFF.
If the battery relay remains energized after the
battery switch has been set to OFF, battery
power will remain on the 28 VDC bus. This will
power the caution and warning panel to allow
illumination of the BATTERY RELAY light,
even if the generator is off.
1-32-K. START ANNUNCIATOR
The START light will be illuminated when the
starter relay is energized. The starter relay will
b e e n e r g i z e d w h e n t h e s ta r t s w i t c h i s
positioned to START as follows:
With the FADEC MODE switch positioned
to A UTO , the s ta rte r rela y w ill s ta y
engaged until the gas producer (N G )
speed reaches 50 ±1%.
With the FADEC MODE switch positioned
to MANUAL, the starter relay will stay
BHT-407-MD-1
engaged until the start switch is released
from the START position.
1-32-L. GEN FAIL ANNUNCIATOR
The GEN FAIL light is controlled by generator
relay. The light will be ON when the generator
relay is de-energized and not connecting the
generator output to the DC bus.
The generator relay will be energized by the
generator control unit/voltage regulator when
generator output climbs through a threshold
of 24 ±2.4 VDC. Prior to the generator relay
being energized, the GEN FAIL light will be
ON. Once the generator relay is energized, the
GEN FAIL light will be OFF.
1-33. PITOT STATIC SYSTEM
The pitot-static system (Figure 1-26) utilizes a
conventional impact air and ambient air
pressure sensing system.
The pitot-static system consists of a heated
pitot tube and right and left heated static ports
and interconnecting tubing.
The pitot tube supplies impact air pressure to
the airspeed indicator and PRCU (paragraph
1-30-D). The two static ports are connected
together to equalize the static pressure, which
is supplied to the airspeed indicator, altimeter,
vertical speed indicator, and PRCU.
1-34. BASIC FLIGHT INSTRUMENTS
The basic set of flight instruments includes an
airspeed indicator, pressure altimeter, vertical
speed indicator, and inclinometer.
1-34-A. AIRSPEED INDICATOR
The airspeed indicator presents airspeed from
0 to 150 knots. The indicator is scaled in 20
knot increments from 0 to 20 knots and in 5
knot increments from 20 to 150 knots. A
maximum speed red line is located at 140
knots and a maximum autorotation speed red/
white line is located at 100 knots.
31 JAN 2007—Rev. 2———1-97
BHT-407-MD-1
MANUFACTURER’S DATA
2
1
3
6
4
7
5
1.
2.
3.
4.
5.
6.
7.
Altimeter
Vertical speed indicator
Airspeed indicator
Pedal restrictor control unit (PRCU)
Pitot static drains
Static ports (heated)
Pitot tube (heated)
6
407_MD_01_0025
Figure 1-26. Pitot Static System
1-98———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
BHT-407-MD-1
1-34-B. ALTIMETER
1-34-G. DIRECTIONAL GYRO
The pressure altimeter presents an altitude
reading in feet above mean sea level (MSL)
based on the relationship between the static
air pressure and the barometric setting on the
altimeter. The barometric setting may be
adjusted to reflect the current barometric
pressure corrected to sea level in inches of
mercury or millibars.
The DG indicator is electrically powered. The
DG is not slaved to a magnetic flux valve and
therefore, must be set prior to takeoff and
periodically during flight to correct for
magnetic variation and gyroscopic
precession. A fail flag will be visible when
power to the indicator is lost warning
directional information is unreliable.
1-34-C. VERTICAL SPEED INDICATOR
(VSI)
The vertical speed indicator presents rate of
climb or descent from 0 to 4000 feet per
minute. A maximum rate of climb red line is
located at 2000 FPM.
1-34-H. TURN-AND-SLIP INDICATOR
T he tu rn -a n d -s lip in d ic at or is po w e re d
electrically and has a pointer that indicates
direction and rate of turn and an inclinometer
that indicates a coordinated or skidding or
slipping turn. A red fail flag will be visible
when power to the indicator is lost, warning
turn information is unreliable.
1-34-D. INCLINOMETER
1-34-I.
The inclinometer consists of a curved glass
tu b e , b a l l, a n d d a m p i n g f lu i d . T h e b a ll
indicates the directional balance of the
helicopter. If the helicopter is in a slip or a
skid, the ball will move off center.
The ENCODING altimeter presents an altitude
reading in feet above mean sea level (MSL)
based on the relationship between the static
air pressure and the barometric setting on the
altimeter. The barometric setting may be
adjusted to reflect the current barometric
pressure corrected to sea level in inches of
mercury or millibars.
1-34-E. OPTIONAL FLIGHT
INSTRUMENTS
Optional flight instruments include an attitude
indicator, directional gyro, turn-and-slip
indicator and encoding altimeter.
1-34-F. ATTITUDE INDICATOR
The attitude indicator is powered electrically
and presents pitch and roll attitudes of the
helicopter in relation to the horizon. A fail flag
will be visible when power to the indicator is
lost warning attitude information is unreliable.
The manual caging knob should be pulled out
to erect the gyro prior to positioning the ATT
switch to ON. The pitch trim knob is used to
adjust the helicopter symbol vertically.
ENCODING ALTIMETER
Additionally, the encoding altimeter provides
coded pulses to the transponder and GPS (if
installed) for altitude reporting.
1-35. NAVIGATION SYSTEMS AND
INSTRUMENTS
The basic navigation equipment consists of a
magnetic compass.
1-35-A. MAGNETIC COMPASS
The magnetic compass is a standard, non
stabilized, magnetic type instrument mounted
on a support which is attached to the right
side of the forward crew cabin. The compass
is used in conjunction with the compass
correction card.
31 JAN 2007—Rev. 2———1-99
BHT-407-MD-1
1-35-B. OPTIONAL NAVIGATION
EQUIPMENT
Optional navigation equipment currently
available through Bell Helicopter consists of a
horizontal situation indicator (HSI), global
positioning system (GPS), VOR indicator,
automatic direction finder (ADF), and
transponder. Basic operational information is
provided for each installation. For expanded
information, refer to the equipment
manufacturer’s operating manual.
1-35-B-1.
HORIZONTAL SITUATION
INDICATOR (HSI)
The optional Kl-525A horizontal situation
indicator (HSI) (Figure 1-4) provides a display
of the gyro stabilized magnetic heading and
helicopter position relative to the VOR and
lo calize r and glide slo pe be ams . Global
positioning system (GPS) course information,
if applicable, may also be displayed on the
H S I. S w itc h in g b et w ee n N AV o r G P S is
selected through a NAV/GPS switch.
Adjustments on the HSI include a heading
select knob for setting the desired magnetic
heading, and a course select knob for setting
the desired course. The course deviation bar
indicates displacement from the selected
radial, track, or ILS localizer beam. The
glideslope pointer indicates displacement
from the ILS glideslope beam. An ambiguity
pointer indicates whether the selected course
will lead TO or FROM the station. The NAV fail
flag warns the pilot of weak or unreliable VOR/
ILS/GPS navigation signals and the HDG flag
warns of gyrocompass failure.
The Bendix/King KCS 55A Gyromagnetic
compass system is powered from the 28 VDC
bus and provides magnetic heading
information, which is displayed on the HSI.
The s ystem include s a K G-102A slaved
directional gyro, KMT-112A magnetic azimuth
transmitter, KA-51B slaving accessory, and a
Kl-525A HSI indicator.
1-100———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
The KA-51B slaving accessory control panel,
located on the pedestal, provides two toggle
switches and a slaving indicator. The SLAVE/
FREE switch engages the slave gyro mode
when in SLAVE position and places the gyro
in the free mode when in the FREE position.
The CW/CCW switch is a momentary switch
which
provides
clockwise
and
counterclockwise manual slewing when the
system is in the free gyro mode.
A slaving meter on the control panel indicates
the instantaneous error between the compass
card presentation and the signal from the flux
valve.
1-35-B-2.
GLOBAL POSITIONING SYSTEM
(GPS)
GPS is a Departm ent O f D efense (DOD)
operated global coverage, satellite-based
navigation system.
The optional KLN 89B GPS (Figure 1-4) is a
long-range, GPS based RNAV System. It
consists of a panel-mounted receiver/display
unit and one antenna. The system is designed
to supply the Pilot with navigation guidance
and position information in three-dimensions:
latitude, longitude, and altitude.
The NAV/GPS push button on the instrument
panel (if installed) will configure the HSI to
depict VOR/LOCALIZER/GLIDESLOPE
information when selected to NAV and GPS
information when selected to GPS.
Additionally, pressing the caution panel test
button will turn on the NAV/GPS switch lights.
Refer to appropriate FMS and the Bendix/King
(Allied Signal) KLN 89B Pilots Guide for
additional information.
1-35-B-3.
VOR INDICATOR
An optional Kl-208 course deviation indicator
(Figure 1-4) may be installed as a primary
VOR/LOCALIZER indicator.
MANUFACTURER’S DATA
An omni bearing selector (OBS) and TO/
FROM ambiguity pointer are provided. The
NAV fail flag warns of weak or unreliable VOR,
localizer signals, or equipment malfunction.
1-35-B-4.
AUTOMATIC DIRECTION FINDER
(ADF)
The optional automatic direction finder set
(Figure 1-4) includes a KR 87 ADF receiver, Kl
227 ADF indicator, and a KA 44B loop/sense
antenna. The receiver includes active/standby
frequency selection and a flight or elapsed
timer.
The right-hand side of the KR 87 receiver will
display either the standby frequency or the
flight or elapsed timer. If it is in one of the
timer modes, press the FRQ button once to
display the standby frequency. Press the FRQ
button again to exchange the active and
standby frequencies.
Press the FLT/ET button to return to one of
the timer modes. The unit will enter whichever
t im e r m o d e w a s d i s p l a y e d p r e v i o u s t o
pushing the FRQ button. Therefore, if it enters
the ET (elapsed timer) mode, press the FLT/ET
button once more to enter the FLT (flight
timer) mode. The flight timer should reset to
zero if the unit is turned off and turned back
on.
The elapsed timer (ET) has two modes:
count-up and count-down. If the display is in
the FLT mode, press the FLT/ET button once
to display the elapsed timer. When power is
applied, the ET is in the count-up mode
starting at 0. When in the count-up mode, the
timer may be reset to 0 by pressing the
RESET button.
To enter the count-down mode, hold the
RESET button in for approximately 2 seconds
until the ET message begins to flash. The
display may now be set to any time up to 59
minutes and 59 seconds. The timer will
remain in this ET set mode (whenever ET
message is flashing) for 15 seconds after a
number is preset, or until the RESET, FLT/ET,
BHT-407-MD-1
or FRQ button is pressed. The preset number
will remain unchanged, until the RESET
button is pressed, at which time it will begin
to count down. When the time reaches 0, it
will begin counting up from 0 and the display
will flash for 15 seconds (regardless of the
current display). While the elapsed timer is
counting down, pressing the RESET button
will have no effect unless it is held for 2
seconds, putting the timer into the ET set
mode.
Pressing the FLT/ET button will exchange the
two timers in the display, or will cause the last
timer that was displayed to reappear if the
s ta n d b y f r e q u e n c y is b e i n g d i s p l a y e d .
Pressing the FRQ button will cause the
s t a n d b y f r e q u e n c y t o r e a p p e a r, a n d
subsequent actuation will cause the active
and standby frequencies to be exchanged.
1-35-B-5.
ATC TRANSPONDER
The optional KT-70 or KT-76A Transponder
(Figure 1-4) is a radio transmitter and receiver
operating at 1090 MHz in transmit mode and
1030 MHz in receive mode. The equipment is
designed to fulfill the role of airborne beacon
under the requirements of the Air Traffic
Control Radar Beacon System (ATCRBS).
Range and azimuth are determined by the
transponder pulsed return in response to
interrogation from the ground radar site. An
identity code number, selected at the front
panel, is transmitted at a Mode A reply. For
mode C altitude reporting capability, the
transponder must be used in conjunction with
a reporting (encoding) altimeter and operated
in ALT mode. After pressing the IDENT button
when interrogated, the transponder will
transmit a special pulse causing the
associated PIP to bloom on the ATC display.
Either transponder can reply on any of 4096
preselected codes.
1-35-B-6.
VHF NAV/COMM SYSTEM
T h e o p t i o n a l V H F N AV / C O M M s y s t e m
(Figure 1-4) can include a KX-155 NAV/COMM,
31 JAN 2007—Rev. 2———1-101
BHT-407-MD-1
o r K X - 1 6 5 N AV / C O M M a s C O M M 1 a n d
KY196A COMM 2.
The KX-155 and KX-165 are VHF NAV/COMM
transceivers which operate within the
frequency range of 118.00 MHz to 136.975
MHz, in 25 kHz increments (760 channels) for
COMM, and 108.00 MHz to 117.95 MHz, in 50
kHz increments (200 channels), for NAV.
Both the KX-155 and KX-165 NAV/COMM have
two displays: the left-hand display for COMM
frequencies and the right-hand display for
NAV frequencies. The COMM display presents
a USE and STANDBY frequency with a T
appearing between the frequency numbers to
indicate operation in the transmit mode. The
NAV display presents a USE and STANDBY
frequency. For both COMM and NAV displays,
the desired frequency is entered into the
STANDBY window and transferred to the USE
window by depressing the transfer button.
B ot h th e C O M M an d N AV, an d U S E an d
STANDBY frequencies are stored in a NVM on
power down, and will be displayed when the
unit is turned on. If an invalid frequency is
de tec ted in the m e m or y w he n po w er is
a p p l ie d , t h e C O M M U S E a n d S TA N D B Y
windows will display 120.000 and the NAV
USE and STANDBY windows will display
110.00.
In addition, when the smaller NAV frequency
selector knob is pulled on the KX-165, the
radial of the active VOR is displayed in the
STANDBY window.
The KY196A is a VHF COMM transceiver
which operates within the frequency range of
11 8 . 0 0 0 M H z t o 1 3 6 . 9 7 5 M H z i n 2 5 k H z
increments (720 channels). The KY196A
d i s p l a y p r e s e n ts a U S E a n d S TA N D B Y
frequency with a T appearing between the
frequency numbers to indicate operation in
the transmit mode. The desired frequency is
en tere d into th e STA ND BY win do w an d
transferred to the use window by depressing
1-102———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
the transfer button. The COMM USE and
STANDBY frequencies are stored in a NVM on
power down, and will be displayed when the
unit is turned on.
1-36. THREE OR FIVE PLACE
INTERCOMMUNICATION
SYSTEM
Optional intercommunication and audio
distribution is accomplished by one KMA
24H-71 audio control panel (Figure 1-4).
The KMA 24H-71 has separate speaker and
headphone isolation amplifiers which are
powered by separate circuit breakers on the
avionic bus, providing a high degree of audio
integrity. The system is equipped with primary
and secondary headphone amplifiers serving
the pilot and the crew/intercom stations
respectively. This provides for the pilot and/or
copilot to isolate from the intercom system.
H e a dp h o n e o u t p u ts c o n ta i n t h e a u d i o
selected by the PHONE pushbuttons and the
MIC rotary selector switch.
The system requires headsets or helmets with
a 300 ohm impedance rating. Military type
helmets with an 8 ohm impedance rating are
not compatible and may cause damage to the
audio control panel.
The KMA 24H-71 audio panel MIC rotary
selector switch has seven positions which are
normally connected as follows:
POSITION
FUNCTION
EMG
Emergency
1
VHF No. 1
2
VHF No. 2
3
Not Used
4
Not Used
5
Not Used
PA
Aft Cabin Speaker
MANUFACTURER’S DATA
The KMA 24H-71 audio panel mixing push
button switches are normally connected as
follows:
BHT-407-MD-1
audio to all headsets by voice activation from
any position.
1-36-B. ISOLATE MODE
SWITCH
FUNCTION
COMM 1
VHF No. 1
COMM 2
VHF No. 2
COMM 3
Not Used
COMM 4
Not Used
COMM 5
Not Used
NAV 1
VOR/ILS
NAV 2
Not Used
DME
Not Used
MKR
Not Used
ADF
ADF
SPKR AUTO
Audio from SPKR
PULL OUT switch
The ICS mode rotary selector switch, located
on the c e nte r co ns ol e, ha s th re e mod e
positions: NORMAL, ISOLATE, and PRIVATE.
Operation in the three modes is as follows.
1-36-A. NORMAL MODE
With the ICS mode rotary selector switch
positioned to NORMAL, and the KMA 24H-71
au dio pan el M IC ro ta ry s ele ct or sw itc h
positioned to COMM 1 or COMM 2 (if
applicable), all headphones will receive audio
from all ICS inputs and selected audio control
panel mixing push button selections.
To operate in the NORMAL (keyed ICS) mode,
the VOX keying knob on the KMA 24H-71
audio panel is to be turned fully counter
clockwise. This will allow audio to all
headsets when the pilot or copilot cyclic
switch is keyed to the first position, the
copilot foot switch is keyed, or any of the aft
ICS drop chord switches are keyed.
To operate in the NORMAL (Hot MIC) mode,
the VOX keying knob on the audio panel is to
be turned fully clockwise. This will allow
With the ICS mode rotary selector switch
positioned to ISOLATE, and the KMA 24H-71
au di o pan el M IC ro tary s ele ct or sw itc h
positioned to COMM 1 or COMM 2 (if
applicable), the pilot will be isolated from the
intercom. Communication between the copilot
and aft ICS stations will be available and
audio from the pilot shall not be heard in the
copilot, aft cabin speaker, or aft ICS headsets.
Audio control panel mixing push button
selections will only be heard through the pilot
headset.
To operate in the ISOLATE (keyed ICS) mode,
the VOX keying knob on the KMA 24H-71
audio panel is to be turned fully counter
clockwise. This will allow audio between the
copilot and aft ICS positions when the copilot
cyclic switch is keyed to the first position, the
copilot foot switch is keyed, or any of the aft
ICS drop chord switches are keyed.
To operate in the ISOLATE (Hot MIC) mode,
the VOX keying knob on the audio panel is to
be turned fully clockwise. This will allow
a u d i o b e t w e e n t h e c o p i l o t a n d a ft I C S
positions by voice activation.
1-36-C. PRIVATE MODE
With the ICS mode rotary selector switch
positioned to PRIVATE, and the KMA 24H-71
au di o pan el M IC ro tary s ele ct or sw itc h
positioned to COMM 1 or COMM 2 (if
applicable), the pilot and copilot will be
isolated from the aft ICS stations.
Hot MIC (voice activated) audio is
automatically provided to all positions in the
PRIVATE mode regardless of the VOX keying
knob position. In addition, audio from the
audio control panel mixing push button
selections will only be heard by the pilot and
copilot.
31 JAN 2007—Rev. 2———1-103
BHT-407-MD-1
1-36-D. AFT CABIN MODE
Aft cabin speaker audio may be selected by
positioning the KMA 24H-71 audio panel MIC
rotary selector switch to PA. Keying the pilot
or copilot cyclic stick switch to the second
position (radio) will allow audio to be heard in
the aft cabin speaker. This will occur with the
ICS mode rotary selector switch positioned to
NORMAL, ISOLATE, or PRIVATE.
1-37. EMERGENCY
COMMUNICATION SYSTEM
In the event of an intercommunication system
failure, emergency communication is
available by positioning the KMA 24H-71
audio panel MIC rotary selector switch to
E M G. W i t h V H F C O M M 1 o p e r a t i n g a n d
selected to the required frequency, keying the
pilots cyclic stick switch to the second
position (radio) will allow two-way
communication. Audio will only be available
in the pilots headset.
Additionally, the emergency communication
system can be activated with the ICS mode
rotary selector switch positioned to NORMAL,
ISOLATE, or PRIVATE.
1-38. AVIONICS MASTER SWITCH
T h e AV I O N I C S M A S T E R s w i t c h a l l o w s
activation and deactivation of all avionics
components simultaneously. Circuit breakers
that are connected to the avionics master
switch are identified by triangles next to the
circuit breakers on the overhead console. The
main 28 VDC bus powers COMM 1 (VHF 1
N AV / C O M M ) a n d t h e a u d i o pa n e l ( I C S
PHONE).
1-39. MISCELLANEOUS
INSTRUMENTS
Miscellaneous instruments include a
combined outside air temperature, clock, and
voltmeter indicator, and an hourmeter.
1-104———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
1-39-A. OUTSIDE AIR
INDICATOR
TEMPERATURE
The outside air temperature is included in a
multifunction indicator mounted in the upper
left area of the instrument panel. The indicator
also displays voltmeter and clock functions.
A button located on the center top of the
instrument changes the top display between
OAT (Celsius and Fahrenheit) and volts (e.g.
28E). When power is applied to the instrument
the display defaults to the voltmeter reading.
Pushing the button will change the display to
temperature.
The temperature is taken by a probe mounted
outside the helicopter in the lower nose
section.
When all electrical power is turned off, the
outside air temperature display disappears.
1-39-B. HOURMETER
An hourmeter is located on the aft wall of the
battery compartment. It can only be viewed
from the inside of the battery compartment.
The hourmeter is a digital instrument that
registers cumulative time, in hours and
tenths. The hourmeter is powered by the
hourmeter circuit breaker located in the
battery compartment. For the hourmeter to
record time, the hourmeter circuit breaker
must be in, the engine N G must be greater
than 55%, and the helicopter weight must be
off the landing gear (the helicopter must be
in-flight).
1-39-C. VENTILATION AND DEFOG
SYSTEM
A vent and defog system is installed at each
crew station. Each side consists of a plenum
with a vent door, electric blower, windshield
defog nozzle, and a control cable.
Control cables, with knobs, are installed
below either side of the instrument panel.
Pulling the cable opens the exterior vent door
to allow outside ram air to enter the cabin
MANUFACTURER’S DATA
BHT-407-MD-1
through an opening at the bottom of the
plen um. The control cables lock in any
position and are released by pressing the
center button on each knob.
instrument lights are controlled by the INSTR
LT rheostat located in the overhead console.
The caution panel is supplied with 28V in
bright and 15V in dim.
An electrically-driven axial flow blower in
each system provides airflow for ventilation
and defogging when the helicopter is on the
ground or hovering. The blower intake takes
air from the cabin and blows it through the
windshield defog nozzle onto the windshield.
Both blowers are controlled by one DEFOG
switch located in the overhead console.
T h e c a u t i o n /w a r n i n g pa n e l d i m m i n g i s
restricted to one position 15V dimming. To
dim the caution panel, the INSTR LT rheostat
must be positioned between dim and bright.
Momentarily positioning the CAUT LT dim
switch to DIM will dim the caution panel lights
to a fixed dim mode.
1-39-D. LIGHTING SYSTEM
The lighting systems include interior and
exterior lighting. The interior lighting system
includes a cockpit utility light, instrument
panel and associated lighting, and aft cabin
lighting. The exterior lighting system includes
landing, position, and anticollision lighting.
1-39-E. COCKPIT UTILITY LIGHT
A removable utility light is secured in a
bracket on the forward side of the control
tube tunnel between the cockpit seat backs. A
long, spiral wound cord permits use
anywhere in the cockpit.
The light will provide a blue or white light
depending on the setting of the selection
switch. It can be used as a spot or flood light
and the intensity of the light is controlled with
the BRT/DIM control knob. Power to the
cockpit utility light is provided through a
5-amp CKPT LIGHTS circuit breaker.
1-39-F. INSTRUMENT
PANEL
ASSOCIATED LIGHTING
AND
The instrument lights are powered by 28 or 5
V DC . A n IN S TR LIG H T S c irc uit b re ak er
provides 28 VDC directly to certain
instruments and also to a 5 volt power supply.
Some instruments (primarily propulsion
instruments) are lighted by the 5 volt power
supply. Other instruments (primarily flight
instruments) are powered by 28 VDC. All
Caution/warning panel segments FLOAT
TEST, BATTERY HOT, ENGINE OVERSPEED,
ENGINE OUT, and RPM are not dimmable. To
test all of the caution panel segment lamps,
press the CAUTION LT TEST switch. This will
also test the lamps for the FADEC mode
switch and NAV/GPS, and Q UIET M ODE
switch lamps (if installed). The PEDAL STOP
switch must be pressed to test its internal
lamps.
1-39-G. AFT CABIN LIGHTING
The aft cabin lighting system consists of two
individual reading lights. Power is provided to
the lights from a 5-amp CKPT LIGHTS circuit
breaker and controlled through the CABIN/
PASS LT switch and two individual reading
light switches.
With the CABIN/PASS LT switch positioned to
OFF, all cabin lights will be OFF regardless of
the position of the two reading light switches.
With the CABIN/PASS LT switch positioned to
C A B I N LT, a l l c a b i n l i g h ts w i l l b e O N
regardless of the position of the two reading
light switches. Positioning the CABIN/PASS
LT switch to PASS LT will allow control of the
reading lights through the reading light
switches.
1-39-H. LANDING LIGHTS
The landing lights consist of one forward and
one downward facing lamp. Both lamps are
exposed for improved cooling.
31 JAN 2007—Rev. 2———1-105
BHT-407-MD-1
Power is provided to the landing lights
through separate relays which are controlled
by the LDG LIGHTS switch located on the
collective switch box. A 2-amp LDG LT CONT
circuit breaker is used to power the coils of
each control relay and a 25-amp LDG LT
POWER circuit breaker is used to power the
landing lights through the control relays.
Positioning the LDG LIGHTS switch to FWD
will turn on the forward landing light and
positioning the switch to both will turn on
both the forward and downward landing
lights.
1-39-I.
POSITION LIGHTS
The position lights consist of a green light
located on the horizontal stabilizer right
vertical fin, a red light on the horizontal
stabilizer left vertical fin, and a white light
located on the tail. Position lights are also
located on the lower cabin with a green light
on the right side, and a red light on the left
side.
Power is provided to the position lights
through a 5-amp POS LT circuit breaker
switch located on the overhead panel.
1-39-J. ANTICOLLISION LIGHT
The anticollision light consists of a single red
strobe light mounted on top of the vertical fin.
Pow er is prov id ed to the strobe from a
separate power supply, which is controlled by
a 5-amp ANTI COLL LT circuit breaker switch
located on the overhead panel.
MANUFACTURER’S DATA
1-42. FIRST AID KIT
The first aid kit is supplied as loose
equipment.
1-43. POINTER 4000 ELT
The Pointer 4000 ELT installation includes a
4000-10 transmitter installed on the left side of
the pedestal, a remote switch on the right side
of glare shield, and an external antenna
mounted on the forward upper cowl.
T h e P o i n t e r E LT i s a s e l f c o n ta i n e d
emergency transmitter capable of manual or
automatic operation. It is designed to
withstand forced landing and crash
environment conditions. Automatic activation
is accomplished by a deceleration sensing
inertia switch. The inertia switch is designed
to activate when the unit senses longitudinal
inertia forces, as required in TSO-C91A.
To configure the ELT to automatically activate
with the remote switch installed, the Master
switch on the ELT must be set to AUTO and
the remote switch must be set to AUTO. This
will allow the ELT to activate when the inertia
switch senses predetermined deceleration
level.
To override the inertia switch and turn on the
ELT manually, the remote switch must be
positioned to ON. This procedure may be
used during unit testing or if an emergency
situation is imminent and pilot wishes to
activate ELT prior to emergency.
1-41. PORTABLE FIRE
EXTINGUISHER
The RESET position of the remote switch is
u s e d t o d e a c t i v a t e a n d r e a r m t h e E LT
transmitter to AUTO mode after automatic
activation by the inertia switch. Helicopter
power is required for remote reset. In case of
inadvertent activation of ELT transmitter with
helicopter power OFF, turn helicopter power
ON, position remote switch to RESET, and
then back to AUTO.
A portable fire extinguisher is mounted
between the cockpit seat backs.
After a forced landing or helicopter accident,
if helicopter receiver is operable, listen on
1-40. EMERGENCY EQUIPMENT
Emergency equipment includes a portable fire
extinguisher, a first aid kit, and optional
Pointer 4000 ELT kit.
1-106———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
121.5 MHz for ELT transmissions. The range
of ELT varies according to weather and
topography. In general, the swept tone signal
can be heard up to 30 miles by a search
helicopter at 10,000 feet. It is recommended to
stay close to the helicopter to permit easier
spotting by airborne searches.
I t m a y a l s o b e d e s i r a b l e t o u s e E LT
transmitter in the portable mode due to a
broken or disabled whip antenna, severed
antenna coax cable, danger of fire or
explosion, temperature extremes in
helicopter, poor transmitting location, or
water ditching with forced evacuation.
To remove transmitter from helicopter:
1.
Bend switch guard away from unit
Master Switch and place switch in
OFF position.
2.
Disconnect remote antenna coax
cable.
3.
Disconnect remote switch cable.
4.
Remove telescopic antenna from
stowage
clips.
Unlatch
ELT
transmitter hold down strap and
remove unit from bracket.
BHT-407-MD-1
5.
Insert telescopic antenna into ANT
receptacle. Extend antenna fully.
6.
Turn ELT transmitter Master switch
to ON position. DO NOT USE AUTO
POSITION.
Consider factors such as terrain, temperature,
and precipitation when choosing a location
for the ELT transmitter to radiate from. The
b e s t t ra n s m i s s i o n m a y b e o b ta i n e d b y
keeping the antenna vertical and setting
transmitter upright on a metallic surface. If
terrain prohibits good transmission (such as
deep valley or canyon), place the transmitter
on high ground or hold in hand in high place.
As cold temperatures have a direct effect on
ELT transmitter battery life, the life of the
battery pack can be extended by placing the
ELT transmitter inside a jacket or coat to keep
the battery warm. Let antenna extend outside
jacket. Keep all moisture and ice away from
the antenna connection and the remote
connector pins.
E n s u r e E LT t r a n s m i t t e r i s t u r n e d O N
continuously (day and night), until rescue
team appears.
For additional information, refer to the Pointer
Aircraft Emergency Locator Transmitter
Operation and Installation booklet.
31 JAN 2007—Rev. 2———1-107/1-108
MANUFACTURER’S DATA
BHT-407-MD-1
Section 2
HANDLING AND SERVICING
2
TABLE OF CONTENTS
Subject
Ground Handling....................................................................................
Covers and Tie-downs...........................................................................
Cover — Engine Inlet.........................................................................
Cover — Pitot Tube............................................................................
Cover — Engine Exhaust ..................................................................
Cover — Oil Cooler Blower Inlet Duct..............................................
Tie-down — Main Rotor.....................................................................
Tie-down — Tail Rotor.......................................................................
Parking — Normal and Turbulent Conditions (Winds Up to
50 Knots).............................................................................................
Mooring (Winds Above 50 Knots) ....................................................
Fuels........................................................................................................
Fuel System Servicing......................................................................
Oils ..........................................................................................................
Engine Oils .........................................................................................
Engine Oil System Servicing ............................................................
Transmission and Tail Rotor Gearbox Oils .....................................
Transmission and Tail Rotor Gearbox Servicing............................
Oil Change — Different Specification ..........................................
Deleted ............................................................................................
Hydraulic Fluids .....................................................................................
Hydraulic System Servicing..............................................................
Paragraph
Number
Page
Number
2-1 ...........
2-2 ...........
2-2-A .......
2-2-B .......
2-2-C .......
2-2-D .......
2-2-E .......
2-2-F........
2-3
2-3
2-3
2-3
2-3
2-3
2-5
2-5
2-2-G .......
2-2-H .......
2-3 ...........
2-3-A .......
2-4 ...........
2-4-A .......
2-4-B .......
2-4-C .......
2-4-D .......
2-4-D-1 ....
2-4-D-2 ....
2-5 ...........
2-5-A .......
2-6
2-6
2-7
2-7
2-8
2-8
2-8
2-9
2-9
2-10
2-10
2-10
2-10
Figure
Number
Page
Number
2-1 ...........
2-4
FIGURES
Subject
Covers and Tie-downs...........................................................................
31 JAN 2007—Rev. 2———2-1
BHT-407-MD-1
MANUFACTURER’S DATA
TABLES
Subject
Commercial Fuels — ASTM D-1655 (Type A and A-1)........................
Commercial Fuels — ASTM D-6615 (Type B)......................................
Military Fuels..........................................................................................
Engine Oils .............................................................................................
Transmission and Tail Rotor Gearbox Oils.........................................
Hydraulic Fluids — MIL-H-5606 (NATO H-515)....................................
2-2———Rev. 2—31 JAN 2007
Table
Number
Page
Number
2-1...........
2-2...........
2-3...........
2-4...........
2-5...........
2-6...........
2-11
2-12
2-13
2-14
2-16
2-17
MANUFACTURER’S DATA
BHT-407-MD-1
Section 2
HANDLING AND SERVICING
2
2-1. GROUND HANDLING
Ground handling of the helicopter consists of
towing, parking, securing, and mooring.
Model 205 or 206 ground handling wheels are
used for towing. Refer to the BHT-407-MM-2,
Chapter 9 for more detailed ground handling
information.
marked TOP is facing upwards. Push the
engine inlet plug into the engine air inlet.
2-2-B.
COVER — PITOT TUBE
WARNING
THE PITOT TUBE CAN BE HOT.
CAUTION
DO NOT TOW THE HELICOPTER IF
THE GROSS WEIGHT IS MORE THAN
5000 POUNDS (2270 KG) FOR
MODEL 205 GROUND HANDLING
WHEELS, OR 4450 POUNDS (2020
KG) FOR MODEL 206 GROUND
HANDLING WHEELS.
2-2. COVERS AND TIE-DOWNS
Protective covers and tie-downs are furnished
as loose equipment and are used for parking
and mooring of helicopter (Figure 2-1).
Additional equipment such as ropes, cables,
cle vis es , ra m p tie-d ow n s, or de ad ma n
tie-downs will be required during mooring.
2-2-A.
COVER — ENGINE INLET
Engine inlet plug assemblies are red and
flame resistant, and each cover is attached
with a red streamer stenciled in white letters,
REMOVE BEFORE FLIGHT. To install the
engine inlet plug, make sure that the side
Pitot tube cover assembly is red, flame
resistant, and attached with a red streamer
stenciled in white letters, REMOVE BEFORE
FLIGHT. To install pitot tube cover, push it
over the pitot tube. Attach the cord.
2-2-C.
COVER — ENGINE EXHAUST
E ng in e ex h au s t co v er is r ed a nd fla m e
r e s i s ta n t , a n d i n c l u d e s a r e d s tr e a me r
stenciled in white letters, REMOVE BEFORE
FLIGHT. A 1/4 inch diameter elastic tie-cord is
attached to cover for securing to engine
exhaust. To install the engine exhaust cover,
push it over the exhaust tailpipe. Attach the
tie-cord.
2-2-D.
COVER — OIL COOLER BLOWER
INLET DUCT
Oil cooler blower inlet duct plug assemblies
are red and flame resistant, and each cover is
attached with a red streamer stenciled in
white letters, REMOVE BEFORE FLIGHT. To
install the inlet plug, push it into the oil cooler
blower inlet duct.
31 JAN 2007—Rev. 2———2-3
BHT-407-MD-1
MANUFACTURER’S DATA
50.0 IN.
(1270.0 mm)
MAXIMUM
DEFLECTION
60.0 LB (27.2 kg)
MAXIMUM LOAD
60.0 LB (27.2 kg)
MAXIMUM LOAD
DETAIL A
4
6
6
5
DETAIL
C
4
SEE DETAIL
5
D
7
DETAIL
B
SEE DETAIL A
3
8
2
6
SEE DETAIL
7
SEE DETAIL
1
DETAIL
D
1.
2.
3.
4.
5.
6.
7.
8.
C
B
Pitot tube cover assembly
Engine inlet plug assembly
Engine exhaust cover
Tie-down assembly
Sock assembly
Line assembly
Tail rotor tie-down strap
Oil cooler blower inlet duct plugs
407_MD-1_02_0001
Figure 2-1. Covers and Tie-downs
2-4———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
2-2-E.
BHT-407-MD-1
TIE-DOWN — MAIN ROTOR
For each main rotor blade, there is a main
rotor tie-down assembly. Each tie-down
assembly has a sock assembly and a line
assembly. Use these assemblies to attach the
blades to the landing gear crosstubes. The
sock assembly is red and has a red streamer
attached to it stenciled with white letters:
REMOVE BEFORE FLIGHT. The line assembly
is made of 0.19 inch (4.83 mm) diameter nylon
and has a ring and an attached flag. The flag
is stenciled with the letters FWD BLADES or
AFT BLADES.
NOTE
Rings are pre-set to apply the
necessary tension to the forward and
aft main rotor blades.
4.
Attach the snaps of the two line
assemblies to the rings of the two
FWD BLADES sock assemblies.
5.
Install the two AFT BLADES sock
assemblies on the ends of the two aft
main rotor blades.
6.
Put a line assembly around each
outboard end of the crosstube of the
landing gear.
7.
Attach the snaps of the two line
assemblies to the rings of the two
AFT BLADES sock assemblies.
Install the main rotor tie-down assemblies as
follows:
CAUTION
DO NOT CAUSE THE MAIN ROTOR
BLADES TO BEND MORE THAN THE
LIM IT S S H OWN IN FI GU R E 2- 1,
DETAIL A.
NOTE
At the same time that you align the
main rotor blades, align the tail rotor
blades with the vertical fin. This will
make it possible to install the tail
rotor tie-down.
1.
Turn the main rotor blades until there
are two blades aft of the fuselage
station of the main rotor hub. When
you look down at the helicopter, the
four blades make an X over the
vertical center line of the fuselage.
2.
Install the two FWD BLADES sock
assemblies on the ends of the main
rotor blades that are forward of the
fuselage station of the main rotor
hub.
3.
Put a line assembly around each
outboard end of the forward
crosstube of the landing gear.
2-2-F.
TIE-DOWN — TAIL ROTOR
The tail rotor tie-down strap is made of 0.025 x
1.0 x 92.0 inches (0.635 x 25 x 2340 mm) nylon
webbing. It is red and stenciled with white
letters REMOVE BEFORE FLIGHT.
Install the tail rotor tie-down as follows:
CAUTION
DO NOT TIE DOWN TAIL ROTOR TO
EXTENT THAT TAIL ROTOR BLADE
FLEXES. THE APPLIED FORCE OF
THE TIE-DOWN SHOULD PROVIDE A
LIGHT CONTACT BETWEEN THE
TAIL ROTOR YOKE AND THE
FLAPPING STOP.
1.
Turn the main rotor blades until there
are two blades aft of the fuselage
station of the main rotor hub. When
you look down at the helicopter, the
four blades should make an X over
the vertical center line of the
fuselage. Align the tail rotor with the
vertical fin.
31 JAN 2007—Rev. 2———2-5
BHT-407-MD-1
2.
3.
2-2-G.
Leaving enough strap material to
extend around the vertical fin, wrap
the tail rotor tie-down strap around
the upper tail rotor blade once
(Figure 2-1, Detail D).
Position the loose ends of the strap
around the upper half of the vertical
fin and tie the two ends together.
PARKING — NORMAL AND
TU R B U L E N T
CONDITIONS
(WINDS UP TO 50 KNOTS)
When winds are forecast to be light or up to
50 knots, park helicopter pointed in direction
from which you expect the highest winds.
Moor helicopter as follows:
1.
Hover, taxi, or tow helicopter to the
specified parking area.
2.
Remove ground handling gear (if
installed).
3.
Attach the main and tail rotor blade
tie-downs.
4.
Install the engine air inlet plugs, oil
cooler blower inlet duct plugs, pitot
tube, and the engine exhaust covers.
5.
Tighten the friction locks on the flight
controls.
6.
Make sure that all switches are in the
OFF position.
7.
Disconnect the battery.
8.
Close and safety all of the doors,
windows, cowlings, and access
panels.
9.
If helicopter is parked outside in a
heavy dew environment, purge
lubricate all of the control bearings
that are open to the air. Do this once
every 7 days. Make sure no voids
exist that could trap moisture.
2-6———Rev. 2—31 JAN 2007
MANUFACTURER’S DATA
2-2-H.
MOORING (WINDS ABOVE 50
KNOTS)
When winds above 50 knots are forecast, park
helicopter pointed in direction from which you
expect the highest winds.
Moor helicopter as follows:
CAUTION
WHEN WINDS ABOVE 75 KNOTS
ARE FORECAST, PUT THE
HELICOPTER IN A HANGAR OR
MOVE IT TO AN AREA WHERE IT
WILL NOT BE AFFECTED BY THE
WEATHER. FLYING OBJECTS
DURING HIGH WINDS CAN CAUSE
DAMAGE TO THE HELICOPTER.
NOTE
If the correct ramp tie-downs are not
available, park the helicopter on an
unpaved area. Use the dead man
tie-downs. Point the helicopter into
the wind and remove the ground
handling wheels.
1.
Attach helicopter
tie-downs.
to
the
ramp
NOTE
Use a mooring clevis at each of the
three jack fittings. This will let you
use a rope with a larger diameter.
2.
Attach the cable, rope, or
manufactured tie-downs to
helicopter jack fittings.
CAUTION
DO NOT CAUSE THE MAIN ROTOR
BLADES TO BEND MORE THAN 50.0
INCHES (1270 MM), AS SHOWN IN
FIGURE 2-1.
the
the
MANUFACTURER’S DATA
main
BHT-407-MD-1
3.
Attach the
tie-downs.
and
tail rotor
4.
If time and storage space are
available, remove the main rotor
blades and put them in a safe
building.
2-3. FUELS
Fuels conforming to following the commercial
and military specifications are approved:
SPECIFICATION
OAT RANGE
ASTM D-1655, Jet A or
A-1
Above -32°C
(-25°F)
Put all of the red streamers inside an
access door so that they will not flap
in the wind.
ASTM D-6615, Jet B
Any OAT
MIL-DTL-5624,
Grade JP-4 (NATO F-40)
Any OAT
5.
MIL-DTL-5624,
Grade JP-5 (NATO F-44)
Above -32°C
(-25°F)
MIL-DTL-83133,
Grade JP-8 (NATO F-34)
Above -32°C
(-25°F)
NOTE
6.
Install the engine air inlet plugs, oil
cooler blower inlet duct plugs, pitot
tube cover, and engine exhaust
cover.
Tighten friction locks on the flight
controls.
CAUTION
MAKE SURE THAT ALL OF THE
SWITCHES ARE IN THE OFF
POSITION, AND THAT ALL OF THE
CIRCUIT BREAKERS ARE OPEN.
7.
Disconnect the battery.
8.
Close and safety all of the doors,
windows and access panels.
9.
Refuel the helicopter to its maximum
capacity.
Refer to the BHT-407-FM-1 for fuel limitations
and use of anti-icing additive.
Fuel listings (Table 2-1 through Table 2-3) are
provided for convenience of operator. It shall
be the responsibility of the operator and the
fuel supplier to ensure fuel used in helicopter
conforms to one of approved specifications.
Refer to Rolls-Royce Operation and
Maintenance Manual for alternate or
emergency fuels.
2-3-A.
FUEL SYSTEM SERVICING
Total capacity
130.5 U.S. gallons
(493.9 L)
Usable fuel
127.8 U.S. gallons
(483.7 L)
Unusable fuel
2.7 U.S. gallons
(10.2 L)
Undrainable fuel
(included in unusable
fuel)
0.7 U.S. gallon
(2.6 L)
CAUTION
SAFETY OR REMOVE ALL OF THE
EQUIPMENT AND OBJECTS IN THE
AREA. OTHERWISE, THE WIND CAN
BLOW THE OBJECTS AGAINST THE
HELICOPTER AND CAUSE DAMAGE.
10. Safety or remove all of the equipment
and objects in the area.
11. When the winds stop, examine the
helicopter for damage.
Fuel system contains two interconnected
cells that are serviced through a single fuel
port located on right side of helicopter. A
grounding jack is provided near fueling port.
An electric sump drain is located in both the
30 APR 2008—Rev. 4———2-7
BHT-407-MD-1
MANUFACTURER’S DATA
forward and aft tanks. They are activated by
buttons located on the right aft lower side of
fuselage. Battery switch must be ON (or
external power applied) and fuel valve switch
must be OFF to activate sump drains.
NOTE
2-4. OILS
Approved oils and vendors are listed in this
section for convenience of operator.
An appropriate entry shall be made in
helicopter logbook when oil has been added
to engine, transmission, or tail rotor gearbox.
Entry shall show specification and brand
name of oil used to prevent inadvertent
mixing of oils.
CAUTION
DO NOT MIX OILS OF DIFFERENT
SPECIFICATIONS. IF OILS BECOME
MIXED, SYSTEM SHALL BE
DRAINED,
FLUSHED,
AND
REFILLED
WITH
PROPER
SPECIFICATION OIL.
2-4-A.
Engine oils (Table 2-4) shall meet engine
manufacturer's
approval.
Consult
Rolls-Royce Operation and Maintenance
Manual for use of oil brands not listed herein.
ENGINE OILS
C e r ta i n o i l s c o n f o r m i n g t o f o l l o w i n g
specifications are approved for use in engine:
SPECIFICATION
OAT RANGE
MIL-PRF-7808
(NATO O-148)
Any OAT
MIL-PRF-23699
(NATO O-156)
OAT above -40°C
(-40°F)
DOD-PRF-85734
———
OAT above -40°C
(-40°F)
Because of availability, reduced
coking, and better lubricating
qualities at higher temperatures,
qualified MIL-PRF-23699 oils are
preferred by engine manufacturer.
NOTE
Long term use of DOD-PRF-85734 oil
may increase probability of seal
leakage in accessory gearbox.
Refer to the BHT-407-FM-1 for engine oil
limitations.
2-4-B.
ENGINE OIL SYSTEM SERVICING
Capacity: 6.0 U.S. quarts (5.7 L).
Engine oil tank is located under aft fairing,
and access doors are provided for filling and
draining oil tank. A sight glass and filler cap
dip stick are provided to determine quantity of
oil in tank.
NOTE
If helicopter engine has been shut
down for more than 15 minutes,
scavenge oil could have drained into
gearbox. Dry motor run engine for 30
seconds before checking oil level. If
not a ccomplishe d, a false high
engine oil consumption rate
indication or overfilling of oil tank
could result. Do not overfill engine oil
tank.
NOTE
NOTE
As per Rolls-Royce, the preferred
engine oils for MIL-PRF-23699 are
Mobil Jet Oil 254 and Aeroshell 560.
2-8———Rev. 4—30 APR 2008
MIL-PRF-23699 and DOD-PRF-85734
oils are not approved for use in
ambient temperatures below -40°C
(-40°F). When changing to an oil of a
MANUFACTURER’S DATA
BHT-407-MD-1
different specification, system shall
be drained and flushed.
Refer to Rolls-Royce Operation and
Maintenance Manual for servicing
instructions and oil filter change procedures.
2-4-C.
TRANSMISSION AND TAIL
ROTOR GEARBOX OILS
Oils conforming to following specifications
are approved for use in transmission and tail
rotor gearbox (Table 2-5):
SPECIFICATION
OAT RANGE
DOD-PRF-85734
—
OAT above -40°C
(-40°F)
MIL-PRF-7808
(NATO O-148)
Below -18°C
(0°F)
NOTE
It
is
recommended
that
DOD-PRF-85734 oil be used in
transmission and tail rotor gearbox
to m a x i m u m e x te n t a l lo w e d b y
temperature limitations. Refer to the
BHT-407-FM-1 for transmission and
tail rotor gearbox oil limitations.
2-4-D.
TRANSMISSION
AND
TAIL
ROTOR GEARBOX SERVICING
Sight glasses are provided to determine
quantity of oil in transmission and tail rotor
gearbox.
Tail rotor gearbox
capacity
0.33 U.S. quarts
(0.31 L)
NOTE
DOD-PRF-85734 oil is not approved
for use in ambient temperatures
below -40°C (-40°F). When changing
to an oil of a different specification,
system shall be drained and flushed.
When adding oil to the main transmission or
tail rotor gearbox, the identical brand and
specification of oil already in each gearbox
shall be used. However, in circumstances
where emergency top-off or inadvertent
mixing may occur, it is acceptable to use oil
with a different brand name within the same
sp ec ifica tion . N o fu rthe r ac tion w ill be
required until the next scheduled oil change,
provided there is no indication of foggy or
hazy oil appearance in the sight gauge.
If oils of different specifications have been
mixed or a foggy or hazy oil appearance
exists, accomplish the required steps per
paragraph 2-4-D-1.
Refer to the BHT-407-MM-2, Chapter 12 for
detailed procedures for draining oil and
changing filters.
2-4-D-1.
OIL CHANGE —
SPECIFICATION
DIFFERENT
W h e n c h a n g i n g t o a n o il o f a d if f e r e n t
specification, accomplish following steps:
NOTE
Transmission oil may partially drain
into freewheel assembly after shut
down. When checking oil levels,
consider this and slope of helicopter
landing surface. If not considered, a
fa ls e o il q u a n ti ty in d ic a ti o n o r
overfilling of gearbox could result.
Transmission capacity
5.0 U.S. quarts
(4.7 L)
NOTE
Refer to the BHT-407-MM-2, Chapter
12 for the maintenance instructions.
1.
Drain transmission, freewheel unit,
and tail rotor gearbox.
2.
Replace transmission oil filter.
3.
Service transmission with proper
amount of approved oil.
30 APR 2008—Rev. 4———2-9
BHT-407-MD-1
MANUFACTURER’S DATA
4.
Service tail rotor gearbox with proper
amount of approved oil.
for use in hydraulic flight control system and
rotor brake.
5.
Operate helicopter for not less than
30 minutes nor longer than 5 hours.
2-5-A.
6.
Drain transmission, freewheel unit,
and tail rotor gearbox.
7.
Service transmission with proper
amount of approved oil.
8.
Service tail rotor gearbox with proper
amount of approved oil.
9.
During first 100 hours of operation
with new oil, check oil sight glasses
closely for indications of foggy or
hazy appearance. If these indications
occur, repeat step 6 through step 9
until eliminated.
HYDRAULIC SYSTEM SERVICING
Reservoir capacity
1.0 U.S. pint
(0.5 L)
Hydraulic reservoir is located on top of
fuselage, forward of transmission, and under
forward fairing. A sight glass is provided to
determine quantity of hydraulic fluid in
reservoir.
Service hydraulic system as follows:
1.
Open and support top of forward
fairing.
2-5. HYDRAULIC FLUIDS
2.
Hydraulic fluids listed in Table 2-6 conform to
MIL-PRF-5606 (NATO H-515) and are approved
Remove cap and fill reservoir until
sight glass is full of hydraulic fluid.
3.
Secure cap and fairing.
2-10———Rev. 4—30 APR 2008
MANUFACTURER’S DATA
BHT-407-MD-1
Table 2-1: Commercial Fuels — ASTM D-1655 (Jet A and A-1)
FUEL VENDOR
ASTM D-1655, JET A
PRODUCT NAME
ASTM D-1655, JET A-1
PRODUCT NAME
American Oil and Supply
American Jet Fuel Type A
American Jet Fuel Type A-1
ARCO (Atlantic Richfield)
Arcojet A
Arcojet A-1
Boron Oil
Jet A Kerosene
Jet A-1 Kerosene
British-American
B-A Jet Fuel JP-1
British Petroleum
B.P. Jet A
California-Texas
B.P. A.T.K.
Caltex Jet A-1
Chevron
Chevron Jet A-50
Cities Service
Citgo Turbine Type A
Continental
Conoco Jet-50
Conoco Jet-60
Exxon Co. USA
Exxon Turbo Fuel A
Exxon Turbo Fuel A-1
Exxon International
Chevron Jet A-1
Esso Turbo Fuel A-1
Gulf Oil
Gulf Jet A
Gulf Jet A-1
Mobil Oil
Mobil Jet A
Mobil Jet A-1
Phillips Petroleum
Philjet A-50
Pure Oil
Purejet Turbine Fuel Type A
Purejet Turbine Fuel Type A-1
Shell Oil
AeroShell Turbine Fuel 640
AeroShell Turbine Fuel 650
Standard Oil of British
Columbia
Chevron Jet Fuel A-50
Chevron Jet Fuel A-1
Standard Oil of California
Chevron Jet Fuel A-50
Chevron Jet Fuel A-1
Standard Oil of Indiana
American Jet Fuel Type A
American Jet Fuel Type A-1
Standard Oil of Kentucky
Standard Turbine Fuel A-50
Standard Turbine Fuel A-1
Standard Oil of New Jersey
Standard Jet A
Standard Jet A-1
Standard Oil of Ohio
Jet A Kerosene
Jet A-1 Kerosene
Standard Oil of Texas
Chevron Avjet A
Chevron Avjet A-1
Union Oil
76 Turbine Fuel
30 APR 2008—Rev. 4———2-11
BHT-407-MD-1
MANUFACTURER’S DATA
Table 2-2: Commercial Fuels — ASTM D-6615 (Jet B)
FUEL VENDOR
ASTM D-6615
PRODUCT NAME
American Oil and Supply
American JP-4
ARCO (Atlantic Richfield)
Arcojet B
British-American
B-A Jet Fuel JP-4
British Petroleum
B.P. A.T.G.
California-Texas
Caltex Jet B
Chevron
Chevron Jet B
Continental
Conoco JP-4
Exxon Co. USA
Exxon Turbo Fuel 4
Exxon International
Esso Turbo Fuel 4
Gulf Oil
Gulf Jet B
Mobil Oil
Mobil Jet B
Phillips Petroleum
Philjet JP-4
Shell Oil
AeroShell Turbine Fuel JP-4
Standard Oil of California
Chevron Jet Fuel B
Standard Oil of Indiana
American JP-4
Standard Oil of Kentucky
Standard Turbine Fuel B
Standard Oil of New Jersey
Standard Jet B
Standard Oil of Texas
Chevron Jet Fuel B
Texaco
Texaco Avjet B
Union Oil
Union JP-4
2-12———Rev. 4—30 APR 2008
MANUFACTURER’S DATA
BHT-407-MD-1
Table 2-3: Military Fuels
COUNTRY
NATO F-34
(JP-8 TYPE)
NATO F-40
(JP-4 TYPE)
NATO F-44
(JP-5 TYPE)
Belgium
BA-PF-7
BA-PF-2
3-GP-24
3-GP-22
3-GP-24
Canada
Denmark
D. Eng. R.D. 2453
MIL-DTL-5624,
Grade JP-4
France
AIR 3405
AIR 3407
AIR 3404
Germany
VTL-9130-006
VTL-9130-007
VTL-9130-010
Greece
MIL-DTL-5624,
Grade JP-4
Italy
AA-M-C.141
AER-M-C.142
AA-M-C.143
Netherlands
D. Eng. R.D. 2453
MIL-DTL-5624,
Grade JP-4
D. Eng. R.D. 2498
Norway
Portugal
MIL-DTL-5624,
Grade JP-4
AIR 3405
Turkey
MIL-DTL-5624,
Grade JP-4
MIL-DTL-5624,
Grade JP-4
United Kingdom
D. Eng. R.D. 2453
D. Eng. R.D. 2454
D. Eng. R.D. 2498
D. Eng. R.D. 2452
United States
MIL-DTL-83133,
Grade JP-8
MIL-DTL-5624,
Grade JP-4
MIL-DTL-5624,
Grade JP-5
30 APR 2008—Rev. 4———2-13
BHT-407-MD-1
MANUFACTURER’S DATA
Table 2-4: Engine Oils
VENDOR
PRODUCT NAME
SPECIFICATION MIL-PRF-7808 (NATO O-148) (FOR OAT ABOVE -40°C/-40°F)
Air BP
BP Turbo Oil 2389
American Oil and Supply
American PQ Lubricant 6899
Bray Oil
Brayco 880H
Mobil Oil
Mobil Avrex S Turbo 256
Mobil RM-184A
Mobil RM-201A
Stauffer Chemical
Stauffer Jet I
SPECIFICATION MIL-PRF-23699 (NATO O-156) OILS (FOR OAT ABOVE -40°C/-40°F)
Air BP
BP Turbo Oil 2380
American Oil and Supply
American PQ Lubricant 6700
Caltex Petroleum
Caltex RPM Jet Engine Oil 5
Castrol
Brayco 899G
Castrol 205
Chevron International
Chevron Jet Engine Oil 5
Hatco Chemical
Hatcol 3211
Mobil Oil
Mobil Jet Oil II
Mobil Jet Oil 254
Royal Lubricants
Royco Turbine Oil 500
Royco Turbine Oil 560
Shell Oil
AeroShell Turbine Oil 500
AeroShell Turbine Oil 560
Stauffer Chemical
Stauffer Jet II (6924)
SPECIFICATION DOD-PRF-85734 (FOR OAT ABOVE -40°C/-40°F)
Air BP
BP Turbo Oil 25
Royal Lubricants
Royco Turbine Oil 555
Shell International
Aeroshell Turbine Oil 555
2-14———Rev. 4—30 APR 2008
MANUFACTURER’S DATA
BHT-407-MD-1
Table 2-5: Transmission and Tail Rotor Gearbox Oils
VENDOR
PRODUCT NAME
SPECIFICATION MIL-PRF-7808 (NATO O-148) (FOR OAT BELOW -18°C/0°F)
Air BP
BP Turbo Oil 2389
BP Turbo Oil 2391
Burmah-Castrol (UK) Ltd.
Castrol 399
Castrol
Brayco 880
Castrol 399
Hatco Chemical
Hatcol 1278
Hatcol 1280
Hexagon Enterprises
Metrex AF Oil 01, 02, 07
Huls America
AOSyn Jet III
PQ Turbine Oil 4236
PQ Turbine Oil 4706
PQ Turbine Oil 4707
PQ Turbine Oil 8365
PQ Turbine Oil 9900
Mobil Oil
RM-248A
RM-272A
NYCO, S.A.
Turbonycoil 160
Royal Lubricants
Royco 808
Shell International
AeroShell Turbine Oil 308
SPECIFICATION DOD-PRF-85734 (FOR OAT ABOVE -40°C/-40°F)
Air BP
BP Turbo Oil 25
Royal Lubricants
Royco Turbine Oil 555
Shell International
Aeroshell Turbine Oil 555
30 APR 2008—Rev. 4———2-15
BHT-407-MD-1
MANUFACTURER’S DATA
Table 2-6: Hydraulic Fluids — MIL-PRF-5606 (NATO H-515)
VENDOR
PRODUCT NAME
Arpol Petroleum
Arpolair 5606
Castrol
Brayco Micronic 756
Castrol Canada
Castrol Aero HF515
Chevron USA
Chevron Aviation Hydraulic Fluid E
(PED 5597)
Chevron PED 6062
Chevron PED 6063
Chevron PED 6064
Chevron PED 6065
Convoy Oil
Convoy 606
Esso SAF
Esso Fluid Aviation Invarol FJ13
Hexagon Enterprises
Metrex Hydrol 1
Huls America
PQ 4140
PQ 9300
PQ 9301
PQ 9302
PQ 9309
PQ 9310
PQ 9311
Mobil Oil
Mobil Aero HFE
NYCO S.A.
NYCO Hydraunycoil FH51
Rohm & Haas
PA 4394
Royal Lubricants
Royco 756
Shell International
AeroShell 41
Technolube
Technolube FB003
2-16———Rev. 4—30 APR 2008
MANUFACTURER’S DATA
BHT-407-MD-1
Section 3
CO NVERSIO N CHA RTS AND TABLES
3
TABLE OF CONTENTS
Subject
Paragraph
INTRODUCTION .........................................................................
CONVERSION TABLES .............................................................
3-1..................
3-2..................
Page
Number
3-3
3-3
LIST OF TABLES
Page
Number
Title
Table
Number
Celsius to Fahrenheit conversion ............................................
Gallons to liters conversion......................................................
Inches to millimeters conversion .............................................
Feet to meters conversion ........................................................
Pounds to kilograms conversion .............................................
Velocity conversion ...................................................................
Standard atmosphere ................................................................
Barometric pressure conversion..............................................
3-1..................
3-2..................
3-3..................
3-4..................
3-5..................
3-6..................
3-7..................
3-8..................
09 Dec 2002
Reissue
3-4
3-5
3-5
3-6
3-6
3-7
3-8
3-9
3-1/3-2
MANUFACTURER’S DATA
BHT-407-MD-1
Section 3
CO NVERSIO N CHA RTS AND TABLES
3
3-1.
INTRODUCTION
This section contains additional information
which may be useful for operational planning
but which is not required for inclusion in the
Rotorcraft Flight Manual. Additional data may
be developed and included as appropriate.
3-2.
CONVERSION TABLES
The conversion tables (Tables 3-1 through 38) provide useful information to assist in flight
planning and operations.
09 Dec 2002
Reissue
3-3
BHT-407-MD-1
MANUFACTURER’S DATA
Table 3-1. Celsius to Fahrenheit conversion
CELSIUS TO FAHRENHEIT
CONVERSION TABLE
°F = (°C x 1.8) + 32°
°C = (°F – 32°) x .555
C. → F.
-62.2
-56.7
-51.1
-45.6
-40.0
-34.4
-31.7
-28.9
-26.1
-23.3
-20.6
-17.8
-15.0
-12.2
-9.4
-6.7
-3.9
-1.1
1.6
4.4
7.2
10.0
12.8
15.6
18.3
21.1
23.9
26.7
29.4
32.2
35.0
37.8
40.6
43.3
46.1
48.9
51.7
54.4
57.2
60.0
62.8
65.6
68.3
C. ← F.
-80
-70
-60
-50
-40
-30
-25
-20
-15
-10
-5
0
5
10
15
20
25
30
35
40
45
50
55
60
65
70
75
80
85
90
95
100
105
110
115
120
125
130
135
140
145
150
155
-112.0
-94.0
-76.0
-58.0
-40.0
-22.0
-13.0
-4.0
5.0
14.0
23.0
32.0
41.0
50.0
59.0
68.0
77.0
86.0
95.0
104.0
113.0
122.0
131.0
140.0
149.0
158.0
167.0
176.0
185.0
194.0
203.0
212.0
221.0
230.0
239.0
248.0
257.0
266.0
275.0
284.0
293.0
302.0
311.0
C. → F.
71.1
73.9
76.7
79.4
82.2
85.0
87.8
90.6
93.3
96.1
98.9
101.7
104.4
107.2
110.0
112.8
115.6
118.3
121.1
126.7
132.2
137.8
143.3
148.9
154.4
160.0
165.6
171.1
176.7
182.2
187.8
193.3
198.9
204.4
210.0
215.6
221.1
226.7
232.2
237.8
243.3
248.9
254.4
C. ← F.
160
165
170
175
180
185
190
195
200
205
210
215
220
225
230
235
240
245
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
320.0
329.0
338.0
347.0
356.0
365.0
374.0
383.0
392.0
401.0
410.0
419.0
428.0
437.0
446.0
455.0
464.0
473.0
482.0
500.0
518.0
536.0
554.0
572.0
590.0
608.0
626.0
644.0
662.0
680.0
698.0
716.0
734.0
752.0
770.0
788.0
806.0
824.0
842.0
860.0
878.0
896.0
914.0
C. → F.
260.0
265.6
271.1
276.7
282.2
287.8
293.3
298.9
304.4
310.0
315.6
326.7
337.8
348.9
360.0
371.1
382.2
393.3
404.4
415.6
426.7
437.8
454.4
482.2
510.0
537.7
565.5
593.3
621.1
648.8
676.6
704.4
732.2
760.0
787.7
815.5
843.3
871.1
898.8
926.6
954.4
982.2
1010.0
C. ← F.
500
510
520
530
540
550
560
570
580
590
600
620
640
660
680
700
720
740
760
780
800
820
850
900
950
1000
1050
1100
1150
1200
1250
1300
1350
1400
1450
1500
1550
1600
1650
1700
1750
1800
1850
932.0
950.0
968.0
986.0
1004.0
1022.0
1040.0
1058.0
1076.0
1094.0
1112.0
1148.0
1184.0
1220.0
1256.0
1292.0
1328.0
1364.0
1400.0
1436.0
1472.0
1508.0
1562.0
1652.0
1742.0
1832.0
1922.0
2012.0
2102.0
2192.0
2282.0
2372.0
2462.0
2552.0
2642.0
2732.0
2822.0
2912.0
3002.0
3092.0
3182.0
3272.0
3362.0
(TABLE I.D. 910630)
3-4
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
Table 3-2. Gallons to liters conversion
GALLONS TO LITERS CONVERSION TABLE
U.S.
GALLON
10
20
30
40
50
60
70
80
90
100
110
120
130
140
150
160
IMPERIAL
GALLON
LITER
8.33
16.65
24.98
33.31
41.63
49.96
58.28
66.61
74.94
83.26
91.59
99.92
108.24
116.57
124.90
133.22
37.85
75.71
113.56
151.42
189.27
227.13
264.98
302.83
340.69
378.54
416.35
454.20
492.05
529.90
567.75
605.60
U.S.
GALLON
IMPERIAL
GALLON
170
180
190
200
210
220
230
240
250
260
270
280
290
300
310
320
LITER
141.55
149.86
158.20
166.52
174.84
183.18
191.50
199.84
208.14
216.48
224.82
233.14
241.56
249.80
258.12
266.44
643.45
681.30
719.16
757.18
795.03
832.89
870.74
908.60
946.45
984.45
1022.16
1060.01
1097.87
1135.62
1173.47
1211.33
(TABLE I.D. 910628)
Table 3-3. Inches to millimeters conversion
INCHES TO MILLIMETERS
CONVERSION TABLE
Inches
0
10
20
30
40
50
60
70
80
90
100
0
1
2
3
4
5
6
7
8
9
mm
mm
mm
mm
mm
mm
mm
mm
mm
mm
−
254.0
508.0
762.0
1016.0
1270.0
1524.0
1778.0
2032.0
2286.0
2540.0
25.4
279.4
533.4
787.4
1041.4
1295.4
1549.4
1803.4
2057.4
2311.4
2565.4
50.8
304.8
558.8
812.8
1066.8
1320.8
1574.8
1828.8
2082.8
2336.8
2590.8
76.2
330.2
584.2
838.2
1092.2
1346.2
1600.2
1854.2
2108.2
2362.2
2616.2
101.6
355.6
609.6
863.6
1117.6
1371.6
1625.6
1879.6
2133.6
2387.6
2641.6
127.0
381.0
635.0
889.0
1143.0
1397.0
1651.0
1905.0
2159.0
2413.0
2667.0
152.4
406.4
660.4
914.4
1168.4
1422.4
1676.4
1930.4
2184.4
2438.4
2692.4
177.8
431.8
685.8
939.8
1193.8
1447.8
1701.8
1955.8
2209.8
2463.8
2717.8
203.2
457.2
711.2
965.2
1219.2
1473.2
1727.2
1981.2
2235.2
2489.2
2743.2
228.6
482.6
736.6
990.6
1244.6
1498.6
1752.6
2006.6
2260.6
2514.6
2768.6
(TABLE I.D. 910627)
09 Dec 2002
Reissue
3-5
BHT-407-MD-1
MANUFACTURER’S DATA
Table 3-4. Feet to meters conversion
FEET TO METERS
CONVERSION TABLE
Feet
0
10
20
30
40
50
60
70
80
90
100
0
1
2
3
4
5
6
7
8
9
Meters
Meters
Meters
Meters
Meters
Meters
Meters
Meters
Meters
Meters
−
3.048
6.096
9.144
12.192
15.240
18.287
21.335
24.383
27.431
30.479
0.305
3.353
6.401
9.449
12.496
15.544
18.592
21.640
24.688
27.736
30.784
0.610
3.658
6.706
9.753
12.801
15.849
18.897
21.945
24.993
28.041
31.089
0.914
3.962
7.010
10.058
13.106
16.154
19.202
22.250
25.298
28.346
31.394
1.219
4.267
7.315
10.363
13.411
16.459
19.507
22.555
25.602
28.651
31.698
1.524
4.572
7.620
10.668
13.716
16.763
19.811
22.859
25.907
28.955
32.003
1.829
4.877
7.925
10.972
14.020
17.068
20.116
23.164
26.212
29.260
32.308
2.134
5.182
8.229
11.277
14.325
17.373
20.421
23.469
26.517
29.565
32.613
2.438
5.486
8.534
11.582
14.630
17.678
20.726
23.774
26.822
29.870
32.918
2.743
5.791
8.839
11.887
14.935
17.983
21.031
24.070
27.126
30.174
33.222
(TABLE I.D. 910626)
Table 3-5. Pounds to kilograms conversion
POUNDS TO KILOGRAMS
CONVERSION TABLE
0
1
2
3
4
5
6
7
8
9
Pounds
Kilograms
Kilograms
Kilograms
Kilograms
Kilograms
Kilograms
Kilograms
Kilograms
Kilograms
Kilograms
0
10
20
30
40
50
60
70
80
90
100
−
4.536
9.072
13.608
18.144
22.680
27.216
31.751
36.287
40.823
45.359
0.454
4.990
9.525
14.061
18.597
23.133
27.669
32.205
36.741
41.277
45.813
0.907
5.443
9.979
14.515
19.051
23.587
28.123
32.659
37.195
41.730
46.266
1.361
5.897
10.433
14.969
19.504
24.040
28.576
33.112
37.648
42.184
46.720
1.814
6.350
10.886
15.422
19.958
24.494
29.030
33.566
38.102
42.638
47.174
2.268
6.804
11.340
15.876
20.412
24.948
29.484
34.019
38.555
43.091
47.627
2.722
7.257
11.793
16.329
20.865
25.401
29.937
34.473
39.009
43.545
48.081
3.175
7.711
12.247
16.783
21.319
25.855
30.391
34.927
39.463
43.998
48.534
3.629
8.165
12.701
17.237
21.722
26.308
30.844
35.380
39.916
44.453
48.988
4.082
8.618
13.154
17.690
22.226
26.762
31.298
35.834
40.370
44.906
49.442
(TABLE I.D. 910648)
3-6
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
Table 3-6. Velocity conversion
VELOCITY CONVERSION TABLE
KNOTS
5
10
15
20
25
30
35
40
45
50
55
60
65
70
75
80
85
90
95
100
MPH
km/HR
5.8
11.5
17.3
23.0
28.8
34.5
40.3
46.0
51.8
57.5
63.3
69.0
74.8
80.6
86.3
92.1
97.8
103.6
109.3
115.1
9.3
18.5
27.8
37.0
46.3
55.6
64.8
74.1
83.3
92.6
101.9
111.1
120.4
129.6
138.9
148.1
157.4
166.7
175.9
185.2
METERS/SEC
2.6
5.1
7.7
10.3
12.9
15.4
18.0
20.6
23.1
25.7
28.3
30.9
33.4
36.0
38.6
41.2
43.7
46.3
48.9
51.4
KNOTS
105
110
115
120
125
130
135
140
145
150
155
160
165
170
175
180
185
190
195
200
MPH
km/HR
120.8
126.6
132.3
138.1
143.8
149.6
155.4
161.1
166.9
172.6
178.4
184.1
189.9
195.6
201.4
207.1
212.9
218.6
224.4
230.2
194.4
203.7
213.0
222.2
231.5
240.7
250.0
259.3
268.5
277.8
287.0
296.3
305.6
314.8
324.1
333.3
342.6
351.9
361.1
370.4
METERS/SEC
54.0
56.6
59.2
61.7
64.3
66.9
69.4
72.0
74.6
77.2
79.7
82.3
84.9
87.4
90.0
92.6
95.2
97.8
100.3
102.9
(TABLE I.D. 910629)
09 Dec 2002
Reissue
3-7
BHT-407-MD-1
MANUFACTURER’S DATA
Table 3-7. Standard atmosphere
STANDARD ATMOSPHERE TABLE
STANDARD CONDITIONS:
TEMPERATURE
PRESSURE
DENSITY
SPEED OF SOUND
15°C (59°F)
29.921 IN. Hg (2116.216 LB/SQ FT)
0.0023769 SLUGS/CU FT
1116.89 FT/SEC (661.7 KNOTS)
CONVERSION FACTORS:
1 IN. Hg = 70.727 LB/SQ FT
1 IN. Hg = 0.49116 LB/SQ IN.
1 KNOT = 1.151 M.P.H.
1 KNOT = 1.688 FT/SEC
DENSITY
RATIO
σ
1
TEMPERATURE
√σ
°C
°F
0
1000
2000
3000
4000
5000
1.0000
0.9711
0.9428
0.9151
0.8881
0.8617
1.0000
1.0148
1.0299
1.0454
1.0611
1.0773
15.000
13.019
11.038
9.056
7.076
5.094
59.000
55.434
51.868
48.302
44.735
41.169
661.7
659.5
657.2
654.9
652.6
650.3
29.921
28.856
27.821
26.817
25.842
24.896
1.0000
0.9644
0.9298
0.8962
0.8637
0.8320
6000
7000
8000
9000
10,000
0.8359
0.8106
0.7860
0.7620
0.7385
1.0938
1.1107
1.1279
1.1456
1.1637
3.113
1.132
-0.850
-2.831
-4.812
37.603
34.037
30.471
26.905
23.338
648.7
645.6
643.3
640.9
638.6
23.978
23.088
22.225
21.388
20.577
0.8014
0.7716
0.7428
0.7148
0.6877
11,000
12,000
13,000
14,000
15,000
0.7155
0.6932
0.6713
0.6500
0.6292
1.1822
1.2011
1.2205
1.2403
1.2606
-6.793
-8.774
-10.756
-12.737
-14.718
19.772
16.206
12.640
9.074
5.508
636.2
633.9
631.5
629.0
626.6
19.791
19.029
18.292
17.577
16.886
0.6614
0.6360
0.6113
0.5875
0.5643
16,000
17,000
18,000
19,000
20,000
0.6090
0.5892
0.5699
0.5511
0.5328
1.2815
1.3028
1.3246
1.3470
1.3700
-16.699
-18.680
-20.662
-22.643
-24.624
1.941
-1.625
-5.191
-8.757
-12.323
624.2
621.8
619.4
617.0
614.6
16.216
15.569
14.942
14.336
13.750
0.5420
0.5203
0.4994
0.4791
0.4595
21,000
22,000
23,000
24,000
25,000
0.5150
0.4976
0.4806
0.4642
0.4481
1.3935
1.4176
1.4424
1.4678
1.4938
-26.605
-28.587
-30.568
-32.549
-34.530
-15.899
-19.456
-23.022
-26.588
-30.154
612.1
609.6
607.1
604.6
602.1
13.184
12.636
12.107
11.597
11.103
0.4406
0.4223
0.4046
0.3874
0.3711
ALTITUDE
FEET
SPEED OF
SOUND
KNOTS
PRESSURE
IN. Hg
PRESSURE
RATIO
(TABLE I.D. 910646)
3-8
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
Table 3-8. Barometric pressure conversion
INCHES TO MILLIBARS
MERCURY
INCHES
0.00
0.01
0.02
0.03
0.04
0.05
0.06
0.07
0.08
0.09
MILLIBARS
28.0−
−
28.1−
−
28.2−
−
28.3−
−
28.4−
−
948.2
951.6
955.0
958.3
961.7
948.5
951.9
955.3
958.7
962.1
948.9
952.3
955.6
959.0
962.4
949.2
952.6
956.0
959.4
962.8
949.5
952.9
956.3
959.7
963.1
949.9
953.3
956.7
960.0
963.4
950.2
953.6
957.0
960.4
963.8
950.6
953.9
957.3
960.7
964.1
950.9
954.3
957.7
961.1
964.4
951.2
954.6
958.0
961.4
964.8
28.5−
−
28.6−
−
28.7−
−
28.8−
−
28.9−
−
965.1
968.5
971.9
975.3
978.7
965.5
968.8
972.2
975.6
979.0
965.8
969.2
972.6
976.0
979.3
966.1
969.5
972.9
976.3
979.7
966.5
969.9
973.2
976.6
980.0
966.8
970.2
973.6
977.0
980.4
967.2
970.5
973.9
977.3
980.7
967.5
970.9
974.3
977.7
981.0
967.8
971.2
974.6
978.0
981.4
968.2
971.6
974.9
978.3
981.7
29.0−
−
29.1−
−
29.2−
−
29.3−
−
29.4−
−
982.1
985.4
988.8
992.2
995.6
982.4
985.8
989.2
992.6
995.9
982.7
986.1
989.5
992.9
996.3
983.1
986.5
989.8
993.2
996.6
983.4
987.8
990.2
993.6
997.0
983.7
987.1
990.5
993.9
997.3
984.1
987.5
990.9
994.2
997.6
984.4
987.8
991.2
994.6
998.0
984.8
988.2
991.5
994.9
998.3
985.1
988.5
991.9
995.3
998.6
29.5−
−
29.6−
−
29.7−
−
29.8−
−
29.9−
−
999.0
1002.4
1005.8
1009.1
1012.5
999.3
1002.7
1006.1
1009.5
1012.9
999.7
1003.1
1006.4
1009.8
1013.2
1000.0
1003.4
1006.8
1010.2
1013.5
1000.4
1003.7
1007.1
1010.5
1013.9
1000.7
1004.1
1007.5
1010.8
1014.2
1001.0
1004.4
1007.8
1011.2
1014.6
1001.4
1004.7
1008.1
1011.5
1014.9
1001.7
1005.1
1008.5
1011.9
1015.2
1002.0
1005.4
1008.8
1012.2
1015.6
30.0−
−
30.1−
−
30.2−
−
30.3−
−
30.4−
−
1015.9
1019.3
1022.7
1026.1
1029.5
1016.3
1019.6
1023.0
1026.4
1029.8
1016.6
1020.0
1023.4
1026.7
1030.1
1016.9
1020.3
1023.7
1027.1
1030.5
1017.3
1020.7
1024.0
1027.4
1030.8
1017.6
1021.0
1024.4
1027.8
1031.2
1018.0
1021.3
1024.7
1028.1
1031.5
1018.3
1021.7
1025.1
1028.4
1031.8
1018.6
1022.0
1025.4
1028.8
1032.2
1019.0
1022.4
1025.7
1029.1
1032.5
30.5−
−
30.6−
−
30.7−
−
30.8−
−
30.9−
−
1032.9
1036.2
1039.6
1043.0
1046.4
1033.2
1036.6
1040.0
1043.3
1046.7
1033.5
1036.9
1040.3
1043.7
1047.1
1033.9
1037.3
1040.6
1044.0
1047.4
1034.2
1037.6
1041.0
1044.4
1047.8
1034.5
1037.9
1041.3
1044.7
1048.1
1034.9
1038.3
1041.7
1045.0
1048.4
1035.2
1038.6
1042.0
1045.4
1048.8
1035.5
1038.9
1042.3
1045.7
1049.1
1035.9
1039.3
1042.7
1061.1
1049.5
6
7
8
9
MILLIBARS TO INCHES
0
1
2
3
MILLIBARS
940
950
960
970
980
990
1000
1010
1020
1030
1040
1050
4
5
INCHES
27.76
28.05
28.35
28.64
28.94
29.23
29.53
29.83
30.12
30.42
30.71
31.01
27.79
28.08
28.38
28.67
28.97
29.26
29.56
29.85
30.15
30.45
30.74
31.04
27.82
28.11
28.41
28.70
29.00
29.29
29.59
29.88
30.18
30.47
30.77
31.07
27.85
28.14
28.44
28.73
29.03
29.32
29.62
29.91
30.21
30.50
30.80
31.09
27.88
28.17
28.47
28.76
29.06
29.35
29.65
29.94
30.24
30.53
30.83
31.12
27.91
28.20
28.50
28.79
29.09
29.38
29.68
29.97
30.27
30.56
30.86
31.15
27.94
28.23
28.53
28.82
29.12
29.41
29.71
30.00
30.30
30.59
30.89
31.18
27.96
28.26
28.56
28.85
29.15
29.44
29.74
30.03
30.33
30.62
30.92
31.21
27.99
28.29
28.58
28.88
29.18
29.47
29.77
30.06
30.36
30.65
30.95
31.24
28.02
28.32
28.61
28.91
29.21
29.50
29.80
30.09
30.39
30.68
30.98
31.27
(TABLE I.D. 910647)
09 Dec 2002
Reissue
3-9/3-10
MANUFACTURER’S DATA
BHT-407-MD-1
Section 4
EXPANDED PERFORMANCE
TABLE OF CONTENTS
Subject
Paragraph
FUEL FLOW CHARTS ................................................................
4-1..................
Page
Number
4-3
LIST OF FIGURES
Title
Fuel flow .....................................................................................
Figure
Number
Page
Number
4-1..................
09 Dec 2002
Reissue
4-4
4-1/4-2
MANUFACTURER’S DATA
BHT-407-MD-1
Section 4
EXPANDED PERFORMANCE
4-1.
FUEL FLOW CHARTS
The fuel flow charts (Figure 4-1) present fuel
consumption during level flight as a function
of altitude, OAT, airspeed, and gross weight.
These charts are based on estimates and
limited flight test data. They are applicable to
the basic helicopter with all doors installed
and without any optional equipment which
would appreciably affect lift, drag, or power
available.
anti-ice operation on fuel consumption. Also,
fuel consumption may vary between engines
under the same operating conditions. It is
recommended, therefore, that the operator
conduct fuel consumption checks to adjust
the presented data as necessary.
The solid line labeled LRC indicates the
optimum long range cruise airspeed for best
fuel economy. LRC is not depicted where it
would occur above the continuous
transmission limit line.
These data do not include the effects of bleed
air heater, ECU, particle separator purge, or
09 Dec 2002
Reissue
4-3
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = SE A LE VE L
O AT = 15°C
100
120
140
TRUE AIRSPEED - KM /HR
160
180
200
220
240
260
400
95
180
M CP TO RQU E LIM IT
380
90
170
360
160
85
V NE
340
80
LR C
60
55
50
FUEL FLO W - LB/HR
TORQ UE - %
65
45
40
140
300
130
280
52
260
30
25
46
110
42
240
30
35
120
50
M AX
EN DURANCE
220
.5
34
38
GW
FUEL FLO W - KG/HR
320
75
70
150
x
10
0
LB
100
200
90
180
80
160
50
50
60
60
70
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
80
90 100
110
120
INDICATE D AIRS P E E D - KNO TS
130
130
140
150
140
M407_MD-1__FIG_4-1_(1_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 1 of 25)
4-4
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 2000 FT
O AT = 11°C
100
120
140
TRUE AIRSPEED - KM /HR
160
180
200
220
240
260
400
180
380
95
170
M CP TO RQU E LIM IT
360
90
160
85
340
80
150
LRC
65
60
55
50
V NE
140
300
130
280
52
260
35
110
46
42
45
40
120
50
M AX
ENDURANC E
240
.5
FUEL FLO W - KG/HR
TO RQ UE - %
70
FUEL FLO W - LB/HR
320
75
220
34
30
200
38
GW
x
0
10
LB
100
90
30
25
180
80
160
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90 100
110
120
INDICATE D AIRS P E E D - KNO TS
130
140
130
150
140
M407_MD-1__FIG_4-1_(2_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 2 of 25)
09 Dec 2002
Reissue
4-5
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 4000 FT
O AT = 7°C
100
120
140
TRUE AIRSPEED - KM /HR
160
180
200
220
240
260
400
180
380
170
95
M CP TO RQU E LIM IT
360
M IN SPEC ENGINE M CP
90
160
340
85
150
70
65
60
FUEL FLO W - LB/HR
TO RQ UE - %
75
320
140
300
130
280
120
260
52
55
50
40
110
46
220
100
42
38
34
200
30
35
30
.5
50
M AX
ENDURANC E
240
45
V NE
LRC
FUEL FLO W - KG/HR
80
GW
x
10
B
0L
90
180
80
25
160
50
60
70
50
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90 100
110
INDICATE D AIRS P E E D - KNO TS
130
120
140
150
130
M407_MD-1__FIG_4-1_(3_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 3 of 25)
4-6
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 6000 FT
O AT = 3°C
100
120
TRUE AIRSPEED - KM /HR
160
180
200
220
140
240
260
380
170
95
M CP TO RQU E LIM IT
M IN SPEC ENGINE + 4%
360
160
90
M IN SPEC ENGINE M CP
340
85
150
320
75
65
60
55
300
FUE L FLO W - LB/HR
TO RQ UE - %
70
V NE
280
52
240
.5
110
50
M AX
ENDURANC E
46
220
45
35
100
42
38
200
34
30
180
GW
x
0
10
LB
90
80
30
25
130
120
260
50
40
140
LR C
FUE L FLO W - KG /HR
80
160
70
140
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90 100
110
INDICATE D AIRS P E E D - KNO TS
130
140
120
150
130
M407_MD-1__FIG_4-1_(4_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 4 of 25)
09 Dec 2002
Reissue
4-7
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 8000 FT
OA T = -1°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
380
95
170
M CP TO RQU E LIM IT
360
160
M IN SPEC ENGINE + 8%
90
340
M IN SPEC ENGINE + 4%
85
150
M IN SPEC ENGINE M CP
320
80
140
300
65
60
55
50
45
LRC
130
280
120
260
V NE
52
50
240
M AX
ENDURANC E
220
30
25
110
100
46
42
200
38
40
35
.5
FUE L FLO W - KG /HR
70
FUE L FLO W - LB/HR
TO RQ UE - %
75
34
180
30
GW
x
10
90
B
0L
80
160
70
140
50
60
70
50
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90
100
110
INDICATE D AIRS P E E D - KNO TS
130
140
120
150
130
M407_MD-1__FIG_4-1_(5_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 5 of 25)
4-8
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 10,000 FT
OA T = -5°C
100
120
TRUE AIRSPEED - KM /HR
160
180
200
220
140
240
260
360
160
M IN SPEC ENGINE + 12%
90
340
150
M IN SPEC ENGINE + 8%
85
320
M IN SPEC ENGINE + 4%
140
80
M IN SPEC ENGINE M CP
300
75
130
280
60
55
50
45
120
260
240
5
M AX
ENDURANC E
220
110
5
2.
50
V NE
100
46
200
90
42
40
38
180
34
35
30
30
160
GW
FUE L FLO W - KG /HR
65
FUE L FLO W - LB/HR
TO RQ UE - %
70
LRC
x
0
10
LB
80
70
25
140
60
120
50
60
70
40
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
140
110
150
120
M407_MD-1__FIG_4-1_(6_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 6 of 25)
09 Dec 2002
Reissue
4-9
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 12,000 FT
OA T = -9°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
360
160
90
340
150
M IN SPEC ENGINE + 12%
85
320
M IN SPEC ENGINE + 8%
300
M IN SPEC ENGINE + 4%
75
60
55
50
FUE L FLO W - LB/HR
TO RQ UE - %
65
LRC
120
260
110
240
52
M AX
EN DUR ANCE
220
V NE
90
42
180
38
34
160
30
30
25
100
46
200
35
.5
50
45
40
130
M IN SPEC ENGINE M CP
280
70
140
FUE L FLO W - KG /HR
80
GW
x
10
80
B
0L
70
140
60
120
50
40
60
50
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
110
140
150
120
M407_MD-1__FIG_4-1_(7_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 7 of 25)
4-10
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 14,000 FT
O AT = -13°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
340
150
85
320
M IN SPEC ENGINE + 12%
140
80
M IN SPEC ENGINE + 8%
280
M IN SPEC ENGINE + 4%
130
M IN SPEC ENGINE M CP
70
120
55
50
45
40
35
110
240
M AX
ENDURANCE
220
100
50
200
90
46
180
V NE
LRC
42
38
160
34
30
25
.5
60
52
TO RQ UE - %
65
FUE L FLO W - LB/HR
260
30
140
GW
x
10
0L
FUE L FLO W - KG /HR
75
300
80
B
70
60
120
50
100
50
60
40
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
140
150
110
M407_MD-1__FIG_4-1_(8_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 8 of 25)
09 Dec 2002
Reissue
4-11
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 16,000 FT
O AT = -17°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
340
150
320
140
80
70
55
50
45
M IN SPEC ENGINE + 8%
M IN SPEC ENGINE + 4%
30
120
M IN SPEC ENGINE M CP
110
.5
240
100
220
50
200
90
46
180
160
38
30
25
120
V NE
34
140
80
LRC
42
40
35
130
52
60
280
260
FUE L FLO W - LB/HR
TO RQ UE - %
65
M IN SPEC ENGINE + 12%
GW
FUE L FLO W - KG /HR
75
300
70
LB
00
1
x
60
M AX
ENDUR ANCE
50
100
50
60
40
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
INDICATE D AIRS P E E D - KNO TS
130
100
140
150
110
M407_MD-1__FIG_4-1_(9_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 9 of 25)
4-12
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 17,000 FT
O AT = -19°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
340
150
320
140
80
300
M IN SPEC ENGINE + 12%
75
130
280
M IN SPEC ENGINE + 8%
60
55
50
45
120
260
M IN SPEC ENGINE + 4%
240
M IN SPEC ENGINE M CP
.5
52
220
100
50
200
90
46
180
80
42
40
160
70
34
140
30
25
120
LRC
38
35
30
110
FUE L FLO W - KG /HR
TO RQ UE - %
65
FUE L FLO W - LB/HR
70
GW
00
x 1
LB
V NE
60
M AX
EN DU RANCE
50
100
50
60
40
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
130
60
70
80
90
INDICATE D AIRS P E E D - KNO TS
100
140
150
110
M407_MD-1__FIG_4-1_(10_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 10 of 25)
09 Dec 2002
Reissue
4-13
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = SE A LE VE L
O AT = 35°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
400
180
95
M CP TO RQU E LIM IT
M IN SPEC ENGINE + 12%
380
90
M IN SPEC ENGINE + 8%
360
85
170
M IN SPEC ENGINE + 4%
80
160
M IN SPEC ENGINE M CP
340
150
65
60
55
50
45
FUEL FLO W - LB/HR
TORQ UE - %
70
320
V NE
140
300
LRC
130
280
52
46
42
240
30
110
38
40
35
120
50
M AX
EN DURANCE
260
.5
34
220
30
GW
FUEL FLO W - KG/HR
75
x
10
0
LB
100
200
90
25
180
80
160
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
130
70
80
90 100
110
120
INDICATE D AIRS P E E D - KNO TS
140
130
150
140
M407_MD-1__FIG_4-1_(11_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 11 of 25)
4-14
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 2000 FT
O AT = 31°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
400
95
180
380
170
M IN SPEC ENGINE + 12%
90
360
M IN SPEC ENGINE + 8%
85
M IN SPEC ENGINE + 4%
340
80
60
55
140
300
V NE
280
LRC
52
260
240
.5
120
50
M AX
ENDURANC E
50
110
46
42
45
40
130
38
220
34
35
30
200
GW
x
0
10
LB
100
90
30
25
FUEL FLO W - KG/HR
320
FUEL FLO W - LB/HR
TO RQ UE - %
65
150
M IN SPEC ENGINE M CP
75
70
160
180
80
160
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90 100
110
INDICATE D AIRS P E E D - KNO TS
130
140
120
150
130
M407_MD-1__FIG_4-1_(12_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 12 of 25)
09 Dec 2002
Reissue
4-15
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 4000 FT
O AT = 27°C
100
120
140
TRUE AIRSPEED - KM /HR
160
180
200
220
240
260
400
180
380
170
360
160
M IN SPEC ENGINE + 12%
85
340
M IN SPEC ENGINE + 8%
80
320
70
65
60
55
50
45
M IN SPEC ENGINE M CP
FU EL FLOW - LB/HR
TO RQ UE - %
75
30
140
300
130
280
V NE
120
260
52
50
M AX
ENDURANC E
240
.5
LRC
110
46
220
100
42
38
40
35
150
M IN SPEC ENGINE + 4%
FUEL FLOW - KG/H R
90
34
200
30
GW
x
10
0
LB
90
180
80
25
160
50
60
50
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90 100
110
INDICATE D AIRS P E E D - KNO TS
130
140
120
150
130
M407_MD-1__FIG_4-1_(13_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 13 of 25)
4-16
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 6000 FT
O AT = 23°C
100
120
TRUE AIRSPEED - KM /HR
160
180
200
220
140
240
260
380
170
360
90
160
340
80
M IN SPEC ENGINE + 12%
320
M IN SPEC ENGINE + 4%
75
TO RQ UE - %
65
60
55
50
45
40
FUE L FLO W - LB/HR
300
70
25
140
M IN SPEC ENGINE M CP
130
280
120
260
52
M AX
EN DURANCE
240
.5
V NE
110
LRC
50
46
220
100
42
38
200
34
35
30
150
M IN SPEC ENGINE + 8%
FUE L FLO W - KG /HR
85
30
180
GW
x
10
0L
B
90
80
160
70
140
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90
100
110
INDICATE D AIRS P E E D - KNO TS
130
140
120
150
130
M407_MD-1__FIG_4-1_(14_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 14 of 25)
09 Dec 2002
Reissue
4-17
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 8000 FT
O AT = 19°C
100
120
TRUE AIRSPEED - KM /HR
160
180
200
220
140
240
260
380
170
360
160
340
85
150
320
75
65
60
55
50
M IN SPEC ENGINE + 4%
280
35
120
260
240
52
50
M AX
ENDUR AN CE
220
.5
110
V NE
LRC
100
46
42
200
38
34
180
30
30
25
130
M IN SPEC ENGINE M CP
45
40
140
M IN SPEC ENGINE + 8%
300
FUE L FLO W - LB/HR
TO RQ UE - %
70
M IN SPEC ENGINE + 12%
GW
x
FUE L FLO W - KG /HR
80
10
90
B
0L
80
160
70
140
50
60
40
50
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
110
140
150
120
M407_MD-1__FIG_4-1_(15_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 15 of 25)
4-18
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 10,000 FT
O AT = 15°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
360
160
340
150
320
80
140
M IN SPEC ENGINE + 12%
300
M IN SPEC ENGINE + 8%
70
60
55
50
45
FUE L FLO W - LB/HR
TO RQ UE - %
65
280
M IN SPEC ENGINE + 4%
30
120
M IN SPEC ENGINE M CP
260
240
M AX
ENDURANCE
52
110
.5
100
50
220
LRC
46
200
V NE
90
42
40
35
130
FUE L FLO W - KG /HR
75
38
180
34
30
160
GW
x
0
10
LB
80
70
25
140
60
120
50
40
60
50
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
140
110
150
120
M407_MD-1__FIG_4-1_(16_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 16 of 25)
09 Dec 2002
Reissue
4-19
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 12,000 FT
O AT = 11°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
360
160
340
150
320
140
80
300
75
M IN SPEC ENGINE + 12%
280
60
55
M IN SPEC ENGINE + 8%
M IN SPEC ENGINE M C P
110
240
5
5
2.
220
50
45
200
46
40
180
50
25
100
V NE
80
38
34
160
30
140
90
LRC
42
35
30
120
M IN SPEC ENGINE + 4%
260
FUE L FLO W - KG /HR
65
FUE L FLO W - LB/HR
TO RQ UE - %
70
130
GW
x
1
LB
00
70
M AX
ENDURANCE
60
120
50
60
40
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
140
150
110
M407_MD-1__FIG_4-1_(17_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 17 of 25)
4-20
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 14,000 FT
O AT = 7°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
340
150
320
140
300
75
130
280
M IN SPEC ENGINE + 12%
60
55
50
45
40
35
FUE L FLO W - LB/HR
TO RQ UE - %
65
M IN SPEC ENGINE + 4%
240
110
M IN SPEC ENGINE M CP
5 2 .5
220
100
50
200
90
46
180
42
160
38
34
30
25
120
M IN SPEC ENGINE + 8%
260
30
140
M AX
ENDUR ANCE
FUE L FLO W - KG /HR
70
LRC
80
GW
x
10
0
V NE
LB
70
60
120
50
100
50
60
40
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
INDICATE D AIRS P E E D - KNO TS
130
100
140
150
110
M407_MD-1__FIG_4-1_(18_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 18 of 25)
09 Dec 2002
Reissue
4-21
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = 16,000 FT
O AT = 3°C
100
120
140
TRUE AIRSPEED - KM /HR
160
180
200
220
240
260
340
150
320
140
300
75
130
280
70
55
50
45
M IN SPEC ENG INE + 8%
5 2 .5
240
5 0 M IN SPEC ENG INE M CP
220
90
46
180
30
100
200
42
40
35
110
M IN SPEC ENG INE + 4%
34
140
30
25
80
LRC
38
160
GW
00
x 1
LB
FUE L FLO W - KG /HR
60
FUE L FLO W - LB/HR
TO RQ UE - %
65
120
M IN SPEC ENGINE + 12%
260
V NE
70
60
M AX
ENDUR ANCE
120
50
100
50
60
40
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
INDICATE D AIRS P E E D - KNO TS
130
140
100
150
110
M407_MD-1__FIG_4-1_(19_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 19 of 25)
4-22
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
PRES SURE ALTITUDE = SE A LE VE L
O AT = 45°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
400
180
380
170
360
M IN SPEC ENGINE + 12%
80
M IN SPEC ENGINE + 8%
340
75
60
55
50
45
FUEL FLO W - LB/HR
TORQ UE - %
65
M IN SPEC ENGINE M CP
140
300
LRC
280
52
30
130
.5
120
46
42
240
110
38
220
V NE
50
M AX
END UR ANCE
260
40
35
150
M IN SPEC ENGINE + 4%
320
70
160
30
34
GW
FUEL FLO W - KG/HR
85
x
10
0
LB
100
200
90
25
180
80
160
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
130
70
80
90 100
110
120
INDICATE D AIRS P E E D - KNO TS
140
150
130
M407_MD-1__FIG_4-1_(20_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 20 of 25)
09 Dec 2002
Reissue
4-23
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 2000 FT
O AT = 41°C
100
TRUE AIRSPEED - KM /HR
140
160
180
200
220
120
240
260
400
180
380
170
360
85
160
M IN SPEC ENGINE + 12%
340
M IN SPEC ENGINE + 8%
75
65
60
55
50
45
40
FUEL FLO W - LB/HR
TO RQ UE - %
70
320
M IN SPEC ENGINE + 4%
25
140
M IN SPEC ENGINE M CP
300
130
280
LRC
52
260
M AX
EN DURANCE
V NE
.5
120
50
46
240
110
42
38
220
34
35
30
150
FUEL FLO W - KG/HR
80
30
200
GW
x
10
0L
B
100
90
180
80
160
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
70
80
90 100
110
INDICATE D AIRS P E E D - KNO TS
130
120
140
150
130
M407_MD-1__FIG_4-1_(21_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 21 of 25)
4-24
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEA N CO NFIGURATION
ENGINE RPM 100%
G ENERATOR 180 AM PS
ZERO W IND
HEATER O FF
ANTI-IC E O FF
BASIC INLET
PR ESS URE ALTITUDE = 4000 FT
O AT = 37°C
100
120
TRUE A IRSPEED - KM /H R
160
180
200
220
140
240
260
380
170
360
160
85
340
80
320
70
55
50
45
M IN SPEC ENG INE M CP
120
260
52
M AX
ENDURANC E
240
LRC
.5
V NE
110
50
46
220
100
42
38
200
34
30
30
25
130
280
40
35
140
M IN SPEC ENG INE + 4%
300
FUE L FLO W - LB/HR
TO RQ UE - %
60
M IN SPEC ENG INE + 8%
FUE L FLO W - KG /HR
75
65
150
M IN SPEC ENG INE + 12%
GW
x
0
10
LB
90
180
80
160
70
140
50
60
50
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
110
INDICATE D AIRS P E E D - KNO TS
130
140
120
150
130
M407_MD-1__FIG_4-1_(22_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 22 of 25)
09 Dec 2002
Reissue
4-25
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 6000 FT
O AT = 33°C
100
120
TRUE AIRSPEED - KM /HR
160
180
200
220
140
240
260
380
170
360
160
340
150
80
320
M IN SPEC ENGINE + 12%
75
140
M IN SPEC ENGINE + 8%
300
60
55
50
45
M IN SPEC ENGINE + 4%
280
130
M IN SP EC E NGINE M CP
120
260
52
M AX
EN DURANCE
240
.5
46
220
110
LRC
50
V NE
FUE L FLO W - KG /HR
65
FUE L FLO W - LB/HR
TO RQ UE - %
70
100
42
40
38
200
34
35
30
30
180
GW
x
0
10
LB
90
80
25
160
70
140
50
60
50
70
60
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
130
70
80
90
100
110
INDICATE D AIRS P E E D - KNO TS
140
150
120
M407_MD-1__FIG_4-1_(23_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 23 of 25)
4-26
Reissue
09 Dec 2002
MANUFACTURER’S DATA
BHT-407-MD-1
FUEL FLOW VS AIRSPEED
CLEAN CO NFIGURATIO N
ENGINE RPM 100%
GENERATOR 180 AM PS
ZERO W IND
HEATER OFF
A NTI-ICE OFF
BASIC INLET
P RES SURE ALTITUDE = 8000 FT
O AT = 29°C
100
120
TRUE AIRSPEED - KM /HR
160
180
200
220
140
240
260
360
160
340
150
320
140
300
M IN SPEC ENGINE + 12%
70
60
55
50
45
40
35
FUE L FLO W - LB/HR
TO RQ UE - %
65
M IN SPEC ENGINE + 4%
260
120
M IN SPEC ENGINE M CP
240
52
50
M AX
EN DURANCE
220
.5
110
LRC
100
46
V NE
42
200
90
38
34
180
30
30
25
130
M IN SPEC ENGINE + 8%
280
GW
FUE L FLO W - KG /HR
75
x
10
0
LB
80
160
70
140
60
120
50
40
60
50
70
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
130
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
110
140
150
120
M407_MD-1__FIG_4-1_(24_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 24 of 25)
09 Dec 2002
Reissue
4-27
BHT-407-MD-1
MANUFACTURER’S DATA
FUEL FLOW VS AIRSPEED
CLEA N CO NFIGURATION
ENGINE RPM 100%
G ENERATOR 180 AM PS
ZERO W IND
HEATER O FF
ANTI-IC E O FF
BASIC INLET
P RESS URE ALTITUDE = 10,000 FT
O AT = 25°C
100
TRUE A IRSPEED - KM /H R
140
160
180
200
220
120
240
260
360
160
340
150
320
140
300
70
60
55
50
45
FUE L FLO W - LB/HR
TO RQ UE - %
65
M IN SPEC ENG INE + 8%
260
30
M IN SPEC ENG INE M CP
240
52
220
.5
100
LRC
46
34
30
160
90
V NE
42
38
180
140
110
50
200
25
120
M IN SPEC ENG INE + 4%
40
35
130
M IN SPEC ENG INE + 12%
280
GW
x
10
0L
FUE L FLO W - KG /HR
75
B
80
70
M AX
EN DURANCE
60
120
50
40
60
70
50
80
90
100
110
120
TRUE AIRS P E E D - KNO TS
60
70
80
90
100
INDICATE D AIRS P E E D - KNO TS
130
140
110
150
120
M407_MD-1__FIG_4-1_(25_OF_25).EMF
Figure 4-1. Fuel flow (Sheet 25 of 25)
4-28
Reissue
09 Dec 2002