Flight Manual S-76a 6w5z4d

-

fk

Sikorsky A Unlted Technologies Company

Q SIKORSKY AIRCRAF'T CORPORATION 2003

SKORSKY AIRCRAFT CORPORATION 6900 MAIN SFREET

P.0. BOX 9729 STRATIPORJl,CT 06615-9129

BRAZILIAN ROTORCRAFT FLIGHT MANUAL

This Rotorcraft Flight Manual is approved by the FAA on khalf of the Centro Tdchnico Aeroespacial fbr Brazilian ed totorcraft, in accordance with "Regulanuentos Brasileitos de Homo@a@ct Aerm4utid' (RBHA)Part 2 1, Section 21.29 Approvedl by: /

e q m h

d

Date: August 30,2000

Robert G. Mann Manager, Boston Aircraft Cettjfimtion Office,ANE-I 50

This Rotorcraft shall be operated in acmdance with the limitations herein established. .,

FAAAPPROVED

ROTORCRAFT FLIGHT MANUAL REVISION NO. 52 SIKORSKY MODEL S-76A PART 1

Sikorsky A United Technologies Company

O SIKORSKY AIRCRAFT CORPORATION 1978

Approved by:

b

R

Louis R. Musacchio Chief, Engineering and Manufacturing Branch, ANE-2 10

U

Date of Approval: NOVEMBER 2 1, 1978 Date of Revision: May 25,2004

Part 1

S - J I I ~ \8.. ~ , ~ STempwary ~ \ Revisions I,M

1

FLIGHT MANUAL PART 1

LOG OF TEMPORARY REVISIONS

REVISION NO.

1

I INCORPORATED1 SUPERSEDED ISSUE DATE

BY REV. NO.

ISSUE DATE

3/12/02

52

5/25/04

1

INSTRUCTIONS:

1.

RETAIN THIS RECORD IN THE FRONT 'OF THE MANWAL.

2.

ON RECEIPT OF TEMPORARY REVfSIONS, INSERTIPAGES ADJACENT TO PAGES AFFECTED IN1ACCORDANCE WTH FlLING INSTRUCTIONS.

.

THIS lLOG WILL BE REVISED EACH TIME THAT TEMPORARY REVISIONS ARE ISSUED, REVISED, ORINCORPORATED BY A NORMAL REVISION.

LOG OF TEMPORARY REVISIONS FAA APPROVED APRIL 9,2002 Revised May 25,2004

Page 1112

Part 1 List of Effective Paees u

FLIGHT MANUAL PART 1 LIST OF EFFECTIVE PAGES

FAA ROTORCRAFT FLIGHT MANUAL

PAGE A B B- 1 C D E F G H J K L M N P

Q R S T

u v W

X Y

z

AA

AB 1

..

I1

iiA iiB iiC iiD iii iv ivA v vi vii viii 1-1 1-2

DATE 5/25/04 5/25/04 Deleted 11/6/03 11/6/03 11/6/03 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 7/26/82 3/9/83 6/28/84 5/14/85 12/12/86 9/22/88 3/16/94 3/16/94 6/17/97 5/15/98 5/25/04 3/16/94 7110187 6/28/84 3/2/89 3/2/89 3/2/89 3116/94 6/29/84 7110187 7/26/82 1/17/83 1/17/83 ll/l6/8 1 5115/92 5/14/85

REV. NO. 52

FAA APPROVED NOVEMBER 2 1,1978 Revised May 25,2004

Part 1 List of Effective Pages

FLIGHT MANUAL PAGE 1-3

DATE -

1-4 1-4A 1-5 1-6 1-6A 1-7 1-8 1-8A 1-8B 1-8B-1 1-8C 1-8D 1-8E 1-8F 1-9 1-10 1-10A 1-11 1-12 1-12A 1-12B 1-13 1-14 1-15 1-16 1-17 2- 1 2-2 2-2A 2-3 2-4 2-4A 2-4B 2-4C 2-5 2-6 2-6A 2-6B 2-6C 2-7 2-8 2-8A 2-9 2-10 2- 10A FAA APPROVED JUNE 19,1996 Revised May 25,2004

REV. NO.

FLIGHT MANUAL PAGE 2- 10B 2- 1OC 2- 1OD 2- 10E 2-1 1 2-12 2-13 2-14 2- 14A 2- 14B 2-14C 2- 14D 2-14E 2- 14F 2- 14G 2- 14H 2-14J 2-14K 2-14L 2-14M 2- 14N 2-14P 2- l4Q 2- 14R 2-14s 2- 14T 2-14U 2-15 2-16 2-16A 2- 16B 2-16C 2-16D 2-16D-1 2- 16D-2 2-16E 2-16F 2-16F-1 2-16F-2 2- 16G 2-16H 2-165 2- 16K 2-17 2-18 2- 18A

DATE 6117/97 10/19/O 1 6/l 7/97 6117/97 5/14/85 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 8/27/79 6/28/84 9/7/90 5115/92 6/28/84 6117/97 6117/97 2/22/85 2/22/85 2/22/85 1/17/83 1/17/83 1/17/83 1/17/83 1/17/83 1/17/83 Deleted Deleted 1/17/83 1/17/83 1/17/83 6117/97 6117/97 3/16/94 3/16/94

Part 1 List of Effective Pages

REV. NO. 46 50 46 46 33 11 11 11 11 11 11 11 11 11 11 11 11 11 11 11 11 11 30 41 42 30 46 46 32 32 32 26 26 26 26 26 26 26 26 26 26 26 46 46 43 43

FAA APPROVED NOVEMBER 2 1,1978 Revised November 6,2003

Part 1 List of Effective Pages

FLIGHT MANUAL PAGE 2-19 2-20 2-2 1 2-22 3- 1 3-2 3-2A 3-2B 3-2C 3-2D 3-3 3-4 3-4A 3-5 3-6 3-6A 3-6B 3-6C 3-7 3-8 3-8A 3-9 3-10 3- 10A 3- 10B 3-1OC 3- 1OD 3-1 1 3-12 3-12A 3- 12B 3-13 3- 14 3- l4A 3- 15 3-16 3-16A 3-16B 3- 16C 3-17 3-18 3-18A 3-18B 3-19 3-20 3-20A

DATE 5/15/92 7/26/82 3/16/94 Deleted 3116/79 5115/92 5/14/85 3123181 7/26/82 7/26/82 7/26/82 5115/92 3/2/89 3/2/89 7/26/82 5/15/92 8/27/79 5/22/80 2/22/85 6/17/97 3/16/94 5115/92 12/12/86 9/7/90 3/2/89 9/22/82 9/22/82 10/19/01 6/29/84 6/29/84 9/7/90 9/7/90 10114/80 5115/92 3/2/89 5115/92 9/7/90 9/7/90 9/7/90 3/2/89 7110187 9/7/90 7/10/87 3124181 7/26/82 4/12/79

FAA APPROVED JUNE 19, 1996 Revised November 6.2003

REV. NO. 42 24 43 24 7 42 33 18 24 24 24 42 40 40 24 42 11 14 32 46 43 42 36 41 40 25 25 50 31 31 41 41 16 42 40 42 41 41 41 40 37 41 37 19 24 8

FLIGHT MANUAL PAGE 3-20B 3-20C 3-2 1 3-22 3-22A 3-23 3-24 4- 1 4-2 4-3 4-4 4-5 4-6 4-6A 4-7 4-8 4-8A 4-8B 4-9 4-10 4-1 1 4-12 4-13 4-14 4-15 4-16 4-17 4-18 4-19 4-20 4-2 1 4-22 4-23 4-24 4-25

DATE 10/14/80 10/14/80 2/22/85 1/17/83 7/26/82 1/9/79 6/6/80 7/26/82 7/26/82 1/9/79 1/9/79 7/26/82 7/26/82 11/2/81 11/2/81 1/17/83 1/17/83 1/17/83 1/17/83 8/27/79 3/23/81 3/23/81 11/2/81 5/14/85 5/14/85 5/14/85 1/17/83 1/17/83 1/17/83 5/14/85 5/14/85 5/14/85 312318 1 2/2/79 5/14/85

Part 1 List of Effective Pages

REV. NO. 16 16 32 26 24 1 15 24 24 1 1 24 24 22 22 26 26 26 26 11 18 18 22 33 33 33 26 26 26 33 33 33 18 5 33

SUPPLEMENT INDEX 6117/97 6117/97 5/15/98 5/15/98 5115/98 5115/98 5115/98 5/15/98

FAA APPROVED NOVEMBER 2 1,1978 Revised November 6,2003

46 46 47 47 47 47 47 47

Part 1 Log of Revisions

FLIGHT MANUAL PART 1 LOG OF REVISIONS

The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes. -

REVISION NO.

PAGES REVISED

1

i, ii, iiA, iiB, iii, iv, v, vi, Section I, All Pages (1-1 through 1-17 and 1-8A), Deleted 1-5,2- 1, 2-2,2-3,2-5,2-6,2-6A, 2-7,2-8,2-9,2-10A, 2-1 1,2-12,2-13,2-14, 2-15,2-16A, 2-17,2-18, 2-19,3-1,3-2,3-2A, 3-2B 3-2C 3-2D 3-4, 3-5,3-6J-7,3-8,3-9,3-10, 3-lOA, 3-11,3-13,3-14, 3-14A, 3-15,3-16, 3-17,3-18,3-19,3-20, 3-20A, 3-21,3-22,3-23, Section IV, All Pages (4- 1 through 4-24)

2

1-1, 2-15

REMARKS

FAA APPROVED

DATE JAN 9, 1979

c.k

\--a&.

Louis R. Musacchio Chief, Engineering

and Manufacturing Branch, ANE-210

Add Night Operation

JAN 16, 1979

R,k Y-i;, Louis R. Musacchio Chief, Engineering

and Manufacturing Branch, ANE-210

3

i, ii, iii, 1-10, 1-10A, 1-16,2-19,3-6A

Added Fuel Flowmeter

JAN 17, 1979

E. E. Hosking for Louis R. Musacchio Chief. Engineering and Manufacturing Branch, ANE-210

FAA APPROVED NOVEMBER 2 1,1978 Revised July 26, 1982

Pa33 Log of Revisions

PART 1

LOG O F REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION

NO. 4

5

PAGES REVISED

REMAmSs

- 2- 3-4, 1-8,l-8A, 1- 10A, 1- 17, 2- 10A, 3-2D, 3- 14A, 3-15, 3-16,4-8,4-9, 4-19,4-23

Add i n t r d ~ c tion, metric scales, various editorial

Inao,1-1, 1-13,2-1,2-2, 2-3,2-5, 2-16,3-19

Pages modified for Canadian M>T repbements

i, iiA, iv, v,

Add IE.R operation

Feb 1, 1979

changes

1-1,l-4,l-5, 1-8,1-17, 2-1OA,2-11, 2- l5,2-16, 2-19,3-10, 3- IOA, 3- 14.4, 3-1?,3-18, 3-19,624

6

FAA APPROVED

DATE

Lords R. Musacchio Chief, Eagmeering and Mamrfacturmg Branch, ANE-210 Feb 2,1979

3

y~.\w&

----------------------------

LOUIS fL M p s a ~ ~ h i ~ Chief, m e e r i n g

1-1,l-5,2-16 (3-19, W T reqairement deleted)

for Cmadian DOT requirements

and M a x u f a c ~ Banch, ANE-210

2-8,2-8A

Added fuel totalizer

Pagesmodified

Feb 21,1919

.

L a ~ lR s M~sacchio Chief, Engineering and bdanrrfacturing Branch, ANE-210

FAA APPROVED NOVEMBER 21, 1978 Revised m y 26, 1982

Part1

Log of Revisions

-

WGHT MANUAL

PART 1

The revisicm date is shown at the bottom af each revised page. A vertical bar an the outer margin indicates tbe Mest revised portion of each page. -

RENISION NO. 7

REVISED iii, ivy 1-1,l-4,

1-5, 1-8, 1-8A, 1-9, 1-14, 1-15, 1-16, 2-1, 2-2, 2-3,2-6A, 2-6B, 2-8, 2-9,2-lOA, 2-11, 2-15,2-16, 2-16A, 3-1, 3-2, 3-6, 3-6A, 3-6B, 3-14, 3-14A, 3-16, 3-16A, 3-20A, 4 8 , 4 9 , 4ll, 4 1 2 , 414,415, 416, 418, 419, 420, 4 2 1 , 422, 4 2 3

-----------------1-1, 1-5,2-1, 2-2, 2-3, 2-16

--

-

-

FAA REMARKS Added metric values, revised power assurance check, and torque available chrfs. Added ECU single pilot VMC and 8pface interior

DATE

APPROVED

MAR 16, 1979

~

Pages m e e d for C a m d i m DOTrequile merits.

8

-

PAGES

P

d

E. E. Boskmg for Louis R. Musacchi0 Chief, m e e r i n g and Manrdactaring Bmnch, ANE-210

i , v, , , 1-3, 1-5, 1-8, 1-8A, 1-8B, 1-12.4,l-13, 2-164 2-16B, 2-16C.2-17.2-19.

Added external Arm 12, 1979 carg~hook, sliding door,

1-5,l-13,2-16A, 2-l6C, 2-17

Pages m e e d for Can&ian DOTreqrrirements

and Collins avionics

Chief, Engineering and ManBranch, ANE-210

FAA APPROVED NOVEMBER 21, 1978 Revised July 26, 1982

5

hi31 Log of Revisions FUGHT MANUAL

PART 1

The revision date is shown at the battom of each revised page. A vertical bar on the oater margin indicates the Mest revised portion of each page, REVISION

NO.

PAGES REVISEI)

REMARES

DATE

FAA APPROVED

Revised altitude June 21,1979 limits

Louis RMosacchio Chief, Engineering and Branch, ANE-210

FAA APPROVED NOVEMBER 21, 1978 Revised July 26, 1982

Part1

- of Revisions

Log

RtGHl MANUAL

PART 1

LOG OF ELFVISlONS The revision date is s h m at the bottom of each revised page. A vertical bar an the outer margin indicates the latest revised portion of each page. REVISION

PAGES

NO.

BEVISED

11

ii, iiA, iiB, iii,v, Pi, ri& V i i i , 1-2, 1-3, 2-1, 2-3, 2-5,2-6A, 2-8,2-9, 2-10,2-lOA., 2-lOB, 2-lOC, 2-11, 2-12, 2-l3,2-14, 2-144 2-14B, 2-14C, 2-MD, 2.-14E, 2-14F, 2-14G, 2-14H, 2-145, 2-1% 2-14L, 2-14M, 2-14N, 2-14P, 2-15, 2-16, 2-164 2-16B, 2-16C, 2-16D,3-2, 3 - 2 4 3-6B, 3 4 C , 3-18,3-18A, 3-20, 3-20C, 47,4-8A, 48B,49,410, 411,412,414, 415,416,420, 421,422,425

12

1-8, 2-1, 2-3, 3-17, 3-18

DATE

REblZARKS Added EAPS. Changed Cate-

FAA APPRDVED

A x . 27,1979

gory '$AffCDP

from 50 to 40 feet. Added DC overtemperatare, ELT, and in£ligkt door emergency procedures. Allow hover for power assurance check

Louis & ~&.Mh~ su2 ~ ~c h iw0 Chief, Ezgimering and Braaca, ANE-210 .

Added informa- Oct 4, 1979 tim on key door locks

FLAPPROVED NOVEMBER 21, Revised July 26, 1982

1978

Part1 ~ o of g Revisions

FLIGHT MANUAL PART 1

LOG OF REVlgONS The revision date is shown at the battom of each revised page. A vertical bar an the outer margin indicates the Latest revised portion of each page.

FAA

REVISOLU' PAGES NO. REVlSED 13 i v y- 2 3 , -71-8-9, 1-11,2-IOC, 2-1OD,2-15, 2-16,3-10, 3-10A, 4-23, 4-25

REMARI(S

Expaad temperatme limits.

DATE Dec 29, 1979

APPROVED

Reservo jam procedure. Added change to Anti-Ice ON

Limits

&eW-

--

Loois R. m c c h i o Chief, Engineeriag

and Mamdactnripg Branch, ANE-210 Added snow protectian Kit P h 76080-30008

.

-

-

May 22,1980

&t

Louis R Musacchio

-6 Engineeand g Branch, ANE210

Page modi-

fied for casaanDOT reqpirements

FAA APPROVED NOVEMBER 21, 1978 Revises July 26, l982

F%rt 1 Section 1 Log of Revisions

NGHT MANUAL PART 1

LOG OF REVMONS The revision date is shawn at the bottom of each revised page. A verl5cdl bar on the outer indicates the latest revised portion of each page.

REVISION NO.

PAGES RJZVEZD

15

i,ii, iiA,ii%, iiC,iv, v, vi, 1-2,1-3, 1-8,

1-8A,1-8B,

FAA

RxxARKS Added U t y hoist and right hand sliding door open o p

DATE

APPROVED

June 6, 1980

Chi&, Engineering and-M Branch, -210 1-13,2-1,2-2, 2-3, 2-16E

Pages modified for Canadian rnT requirements

FAA APPROVED NOVEMBER 21, 1918 Revised h l y 26, 1982

Part 1 Section 1 Log of Revisions

FLIGHT MANUAL

PART 1 LOG OF REITSIONS

The revision date is shown at the bottom of each revised page. A vertical t3ar on the outer margin indicates the latest revised portion of each page. REVISION

PAGES

NO.

REVISED

17

FAA

DATE

REl\bARIZS

1-2, 1-3, Removed ECU 1-8, 2-lOC, operating Limits due to high 15. 2-15, 2-16, 2-16G, 3-10, 3 -IOA, 3 l6A, Updated perfor3-17, 4-8, mance data with 4-8A, 4-83, engine anti-ice 4-9, 4-11, onand removed 4-l2, 4-14, the 4000 ft alt 4-15, 4-16, limit for lake 4-17, 4-18, ofk and hndings 4-19. 4-20. with engine - anti4-2c 4-22, ice an. 4-23, 4-25

&C

APPROVED

I?, 1980

-

.

atfan of pressure

Chief, Ehgineering andak~f2ctaring

Bsanch, ANE-210

Added snow proteckit P/N 7607630006-012.

18

1-2,l-3, 1-7,3-2, 3-2B,3-20, 3-3,411, 4-12, 4-18, 4-19, 4-23, 4-25

. .

Increased density Mar 23, 1981 nftitnde limft to 11,000 feet

$4-

Revise procectaretof~llorP after m e

Louis R M-chio Chief, h . I p m q d -

-@=

Branch, ANE210

f a p a

own

FAA APPROVED NOVEMBER 21, 1978 Revised July 26, 1982

FLIGHT MANUAL

PART 1

LOG OF REVISIONS The revision date is shown at the bottom af each revised page. A vertical bar on the outer margin mdicates the latest revised portion of each page.

PAGES

REVISIdU-

RE-

NO. 19

iv, ivA. vii, 1-8, 2-6, 2-10B, 2-10C, 2-lQD, 2-15, 2-16, 3-lOA, 3-lOB, 3-lOC, 3-11, 3-18A, 3-19

REMARKS Added AFCS Phase IU and Editorial changes

Added hsb3led Pwer Assmance Check and Power Trend Check

FAA APPROVED

DATE

L,rW ; -L Louis R Musacchio Chief, Engineering and Manufacturing Branch, ANE-210

6/26/81

ELMusacchio

Chief, Engineering and Man-

Branch, AXE-210

21

Added note on No. 2 servo

2-8

9/1/81

Kessare and bhrrufactraiog Branch, ANE-210

1 - 1 , l - 2 , 1-3, 1-4, 1-6, 1-8C, 1-17, 4 6 A , 4-7, 4 8 , 4-84 48B, 4 9 , 413, 414, 4-15, 4 1 6 , 4 2 0 , 4-21, 422, 4 2 5

Ln~zeasipg 3- gmss weight tolO,SOO

P-• and Branch, ANE-210

F M -APPROVED NOVEMBER 21, 1978 Revised July 26, 1982

Partlsectionl Log of Revisions

N G H T MANUAL

PART 1 LOG OF ECEmoNS The revision date is shown at the battom of each r e -page. A vertical bar on the outer margin indicates the latest revised portion of each page.

REVISION

NO. 23

PAGES FtEVISED

iv, ivA, vi, vii, viii, 2-5, 2-10, 2-lOC, 2-lOD, 2-11,2-14% 2-14S,2-16, 2-16B. 2-16F-1

REMAIiICS

FAA APPROVED

DATE

Added: Snow 11/16/81 Protection git with Continoo~s

Ignition, Master

Warning and Caution AdPisory Failme

Emergency Proced-me, In Fliglst Power Aswrarrce

and Hoist lockoatuf ECU and Vent Blower

Chief, Engineering

B m h , ANE-210

FAA APPROVED NOVEMBER 21, 1978 Repised July 26, 1982

Part 1 Section 1

Lw of Revisions

The revision date is shown at the buttom of each revised page. A vertical bar on the rear xiargin indicates the latest revised partion af each page. REVISION NO. 24

PAGES REVISED

DATE

FAA APPROVED

i, ii, iiA, iiB, iiC, iiD, iii, ivyivA, 7/26/82 -V, 1-2,l-3, 1-6A, 1-?,I-8A, 1-83, 1-11,l-14,l-15,l-16, %2A, 2-5, 2-6,2-6A, 2-8,2-9,2-10, 2-IOA, 2-10B, 2-10C, 2-11,2-16B, 2-160, 2-16G, 2 - 1 8 , 2 - l a , 2-20,Z-21, 3-2, 3-2C, %2D, 3-3,3-6,%6A, Chief, Boston Air& 3-10A, 3-10B, 3-10D, 3-12,3-12A, 3-15,3-17, 3-20,3-22, 3-22A,4-1, Certification Branch, 4-2,45,46 ANE-150 ,

REMARE=S: Added Note on Smake Detector Baggage Warniag Ligkt Droop C o q x m a b r Amplifier PrefEgkt Check AFCS Oscillatory ldaBmd5on Alternate Airspeed hdicator

Deleted

Note on No. 2 HpdraaZic Pressure and St-g Pages 2 4 4 and 2-22

&@nes,

Revised Table of Contents Color hlshments (3fLgures) -or Check rJ1 me speed Fire Detector Test O p e M o ~ Fuel Crossfeed after Zagbe Failure Procedure Tail Rotot Control System Failure Fixed Pitcb Setbing Paerndl Cargo Hook PreWM Fnnciioml Check Snow Protection Kit Airspeed Calibration Figares Tg Check in PrEagineOilSystemFaiIPre

FAA APPROVED NOVEMBER 21, 1978 Revised M y 26, 1982

PartlSed5onl

Log af Revisions

RlGHl MANUAL PART 1

The revision date is shown at the bottom af each

revised page. A vertical har on the rear margin indicates the fatest revised portion af each page. REVISON

NO. 23

PAGES REVISED

, 1- 5, 1-8, 2- l4R, 3-10A,3-103, 3-10C,3-IOD

1 ,

RElKmKS

Removal af Pitch Bias Actuator from Operating Limitations

DATE

FAA APPROVED

9/22/82

Arthur Pidgeon Chief, Boston Aircraft

Certification Branch, ANE- 150

Added RHDR light Revised engine start NOTE, and Tg

1 /17/83

-=

Ftevised and corrected art

and text. Added nitrogen substitute for helium in flo-

tation system. A m J. Pidgeon Manager, Boston Aircraft Certification B m h , ANE-150

27

1-8,2-12-18,2-21, 3-10A

Revise fnnction af .3/9/83 the pitch bias artaator.

FAA APPROVED NOVEMBER 21, 1978 Revised MARCE 9, 1983

Part 1 Section 1 Log of Revisions

FUGCrr MANUAL

LOG OF RFVISIONS

The revision date is shown at the battam of each revised page. A vertical bar on the rear X Y X W ~ ~ indicates the Latest revised portion of each page.

REVISION

PAGES

NO.

REVISED

29

2-6, 2- 10B

REMARKS

Revised to provide for a p t i d

DATE

APPROVED

11/3/83

PBA

Aircraft celw5cajian Brapch, ANE-150

30

iiA, ZB, iv, 1- 15, 2-1,2-2, 2 - 2 4 2-3, 2-4,2-6, 2-6.4 2-6B, 2-8, 2-u, 2- 146, 2- 141, 2-164 3-30, 3- lOA,

Revised engine 6/28/84 starting information. servo unit jam g ~ ~ takeoff and approach information, oil prelimits. rotor brake checks, efectrical fire peedares and edit0ria.l corrertions.

,

Manager, Boston Aircraft Certjfietim Branch, ANE-150

3- 15

FAA APPROVED NOVEMBER 21, 1978 Revised JUNE 28, 1984

LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on tbe rear margin indicates the Mest revised portion of each page.

'REVISION

PAGES

NO.

REVISED

31

iv, ivA, 3-12 Revised tail rotor 6/29/84 3- 1% 3- 13 control sPstem u e s Gith the self-centering

FAA

l3EMARKS

APPROVED

DATE

tail rotor quadrant

Manager, Boston AirCmft certifica-

tion Office, ANG 150

32

ii, 1-11, Added mnmrnl 2/22/85 1 -- 3 , rotor btake. Revised £light 2-2,2-2A, 2-6,2-6~, control system checks, contirmoas 2-6B,2-8, 2-842-lOC, ignition usage, fwl priming, chip Manager. Boston 2-16,2-16A, 2- 163,2-18A, detectors, and &reraft -Cert&eation Office, ANE2-19, 3-7, dik3-9, 3-21 150,

-

Increased maxigross

weight b 10.500

Po-.

5/14/85

&#.U

Arthur J. Pidgeon Manager. Boston Aircraft, Certification Office, ANG 150

FAA APPROVED NOVEMBER 21, 1978 Revised MAY 14, 1985

Part 1 Section I Log of Revisions

A~~G?BB~~Y//&. FUGW MANUAL PART 1

LOG OF REVISIONS T h e revision d a t e is shown at t h e bottom of each revised page. A vertical bar on the rear margin indicates t h e latest revised portion of each page.

REVISION NO. 34

35

PAGES REVISED

1-6A,2-1, 2-2, 2-3, 2-4, 2-4A, 2-43, 2-6A, 3- 12B, 3- 13

3- 8

REMARKS

DATE

APPROVED

Revisedfire 3/17/86 warninglights, preflight external check, and Manager, Boston Aircraft Certification airspeed indicator Office, ANE-150 Revised fuel crossfeed limits

3/18/ 86

u

x

=d

Manager, Boston Aircraft Certification Office, ANE-150

36

3- 10-3- 10A Revised flight control s e r v o emergency

12112/86

procedures .

Aircraft Certification Office, ANE: 150

FAA APPROVED NOVEMBER 21, 1978 Revised DECEhlBER 12. 1986

Part 1 h g of Revisions

FLIGHT MANUAL

PART 1 LOG OF REVISIONS The revision date is shown at the botxom of each revised pqe. A vertical bar on the outer rnar*~indicates the latest revised portion of each page.

'REVISION PAGES NO. REVlSED

REMARKS

DATE

FAA APPROVED

Revised fuel cross- 7110187 feed limitations, AC generator test procedure, hoist duty cycle, engine priming and emergency procedures for low oil pressure, battery overtemperame and dual dc generator failure. - -

-

~

M a q e r , Boston Aircraft Certification Office, A??-150

38

1-3, 1 4

Revised category 'B" Maximum Takeoff Gross wei* am.

4/13/88

Manager, Boston Aircraft Certification Office, ANE-150

39

2-1, 2-2, 2-2A, 24A,

Revised Normal Procedures.

9/22/88

Manager, Boston &craft Certification ME,ANE-150

FAA APPROVED NOVEMBER 21,1978 Revised September 22, 1988

Part 1 of Revisions

J&

PART 1 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page.

REVISION NO. 40

PAGES REVISED

FEMARKS

s3,iic, a, Revised transmis1-9, 1-11, 1-12, 1-12A, 1-12B, 1-13, 2-1, 2-2, 26C, 2-7, 2-SA, 2-9, 2-16K, 2-17, 34A, 3-5, 3 8 , 3-lOA, 315, 216A, 316B, 3-17, 3-18A

FAA APPROVED

DATE 3/2/89

sion torque Limit, placards, main gear box oil level, engine oil reservoirs, and flight controls and servo systems checks, hoist duty cycle, N2 runaway, fli@t control damper jam, and master warning and caution advisory fail"

electrical system failures.

Manager, Boston

Supplement Index

Aircraft Certification Office, ANE- 150

Revised landing gear, f k , and semo damper jam ;-rp O F ating limits;and xlOmd

9/7/90

F-

dures; added JF-8 as approved fuel.

& d f d

Supplement Index

Manager, Boston Air& -cation Office, ANE- 150

FAA APPROVED NOVEMBER 21, 1978 Revised March 16, 1994

b

1

Log of Revisions FUGHTMANUAL

PART 1 LOG OF REVISIONS The revision date is shown at rhe boaom of each revised page. A vertical bar on the outer mar_ein indicafes the larest revised @on of e s h page.

'REVISION PAGES NO.

42

REVISED 1-1, 1 4 , I-SA, 1-12, 2-6A, 2-9, 2-IOA, 2-la, 2-19, 3-2, 3-4, M A , 3-8A7 3-9, 3-14A, 3-16

FAA

F?EMAMS

DATE

Revised opaaring

5/15/92

APPROVED

fimirarilms.

Delefed elecaonic N2 overspeed

Jv

-

-M=%a7 Aircraft Cereification Office, Am-150

43

1 2-18, 2-18A, 2-21,3-8, 3-8A

Claxifiedkxggage

3/16/94

flow M

1 e . Flotation System Te and Main Gear Box Chip-

E==w=Y

FAA APPROVED NOVEMBER 21, 1978 Revised March 16, 1%

Part 1 Lo,o of Revisions

FLIGHT MANUAL

PART 1 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each pase. A vertical bar next to the page number indicates the page has been retyped with no significant changes.

REVISION NO. 44

PAGES REVISED

REMARKS

DATE

1-12, 3-8

Revised transrnission oil spec; main gear box Iow oil pressure emergency procedure.

9112/95

3, 4,5, 6

Updated Supplement Index

-

&> 45

1-8B-1, 1-8C Revised external cargo hook Iimitations

FAA APPROVED

Manager, Boston Aircraft Certification Office, ANE-150 3/ 19/96

Index 1 thru 7 Updated Supplement Index - /)-

~ k a g e r Boston , Aircraft Certification Office. ANE-150

FAA APPROVED NOVEMBER 21, 1978 Revised June 17, 1997

Part 1 Log of Revisions

FLIGHT MANUAL PART I

LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes.

REVISION NO.

PAGES REVISED

REMARKS

46

1-4,2-1, 2-4A, 2-4B, 2-4C, 2-6C, 2- 1OA, 2- 1OB, 2- lOC, 2- 1OD, 2- 1OE, 2-14U, 2-15, 2- 16K, 2- 17, 3-8

Revised Category A and B operation, revised Exterior Check procedure instruction, added Warning for aft engine cowl latches and straps, added air conditioner Note, added AFCS Check Note, and added Approach and Landing Note.

Index 1, 2, 3,

Updated Supplement Index

4, 5,6, 7

47

1-8D

Revised Minimum Flight Crew requirements for Utility Hoist operations.

DATE

FAA APPROVED

6/17/97

S6r I

Manager, Boston Aircraft Certification Office, ANE- 150 5'1 5198

-Fbr

FAA APPROVED NOVEMBER 2 1,1978 Revised May 15, 1998

/

Part 1 Log of Revisions

FLIGHT MANUAL PART I LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes.

REVISION NO. 48

49

50

51

PAGES REVISED 2-6, 2-8

2-2,2-4A

1-5,2- 10C, 3-1 1

1-5

REMARKS Revised Rotor Brake Procedures contained in Normal Procedures. Section 11.

Revised Exterior Check - Main Rotor Blades.

Revised Minimum Flight Crew Criteria, Taxiing procedure, and pitch bias actuator reference. Change IFR equipment requirements

DATE

FAA APPROVED

7115/99

Manager, Boston Aircraft Certification Office, ANE- 150 9/23/99

Manager, Boston Aircraft Certification Office, ANE- 150 10119101

&Manager, Boston Aircraft CertificationOffice, ANE-150

1 1/6/03

Certification Office, AN€-I50

52

2- 1,2-3,2-4

Add cyclic base inspection, Incorporate horizontal stabilizer inspection (Temporary Revision No. I), add tiedown ring check.

5/25/04

Manager, Boston A Certification Office, AN€-I50

FAA APPROVED NOVEMBER 2 1,1978 Revised May 25,2004

REVISKIN NO.

PAGES REVISED

CfA

REMARKS

CTA APPR-

DATE

AUGUST 30,2000

Reviead: M Q v e m k 6.m

APPROVED

Part 1 Table of Contents FUGHT MANUAL

PART 1 TABLE OF CONTENTS INTRODUCTION ........ (BACK OF lTI'LE PAGE .CAN DOT ONLY) LIST OF EFFECTIVE PAGES ........................ A LOG OF REVISIONS PAGE ......................... F SErnON I OPERATING LIMITATIONS ...............

PAGE 1-1

WEIGHT LIMITS .........................

1-1

CATEGORY "A" OPERATIONS .......... CATEGORY "B" OPERATIONS ..........

1-1 1-1

CG (CENTEROF GRAVITY) LZMITS .......

1-1

.................... . ..

1-4

LOADING LIMITS

TYPES OF OPERATION ...................

1-4

CATEGORY "A" AND "B" .............. CATEGORY "B7'ROTORCRAFT . EXTEWNAL LOAD COMBINATIONS ...

1-5

MINIMUM FLIGHT CREW ................

1-5

AIRSPEED UMtTS .......................

1-6a

ALTlTUDE L l h a S .......................

1-7

AMBIENT TEMPERATURE L I h a S ........

1-7

FLIGHTJAMITS .........................

1-7

1-4

SLIDING CABIN DOOR L;LMITS ............

1-8A

FLOTATION SYSTEM LMlTS .............

1-8a

EXTERNAL CARGO HOOK L I h a S ........

1-8B

WEIGHT LIMITS ...................... AIRSPEED LIMITS .....................

1-8B 1-8B

lJTm7-YHOIST LIMITS ..................

1-8C

WEIGHTUMITS ...................... MINIMUM FLIGrn CREW .............. AIRSPEED LIMITS .....................

1-8D 1-8D 1-8D

FAA APPROVED NOVEMBER 21. 1978 Revised March 16. 1994

I

Part1 Table of Contents

TABLE OF

CONTENTS (Cod)

SEC'JTON

PAGE

........................... 1-8E FUEL .................................. I-8E O IL ................................... 1-10 OIL TEWERATORE I;IMITS ............. 1-10 OIL PRESSURE L l M r B ................. 1-10

~ ~ L I M Z T S

FDEL PRESSURE LIMITS (IF FUEL PRESSURE GAGES ARE INSTALLED) FDEL FLOW L;IMITS (B?FUEL FLOW-

.... 1-10 METERS ARE INSTALLED) -.--..-...... 1-10 ENGINE TORQ'UE LIMITS ............... 1-10 (GAS PRODUCER) S P E W LIMITS ....- 1 - 10A -NN1 :, (POWERTURBINE) SPEED

-L;IMITS TR -

-1-10A ............................. TURBINE) INLET

................

TEMPERA= L;TMITS 1-11 steady State Limits ................... 1-11 Startiag a Shntdo~n.................. 1-11 Power Transients ................... 1-11

.......................

TORQUE I3MrrS 1-11 Dual-Engine Operation ................ 1-11 Single-Engine operation ............... 1-11

OIL .................................... . oE TEMPERATURE LIMITS ............. ~

OIL PREssuRE Z;IMITS

ROTOR STOPPED ROTOR TURNING

~

.................

1-11 I- ii 1-12

..-....................1-12

....................... 1-12

. ...............................

ROTOR BLADE L;LMITS ROTOR STOPPED

1-12

............. 1-12 ............ 1.1%

DUALENOP~TION SINGLE-ENOPERATION ENGIhF COWLING REMOVED OR E I N G D BACK

..................... 1.12A

...............................1-12A CARGO BOOK/HOET LIMIT'S ............. 1-13

PLACARDS

FAA APPROVED NOVEMBER 21. 1978 Revised J u l y 10. 1987

Part I Table of Contents FLIGHT MANUAL

TABLE OF CONTENTS (Cont) PAGE

SECTIOh'

11

NORMAL PROCEDURES

....................

2-1

NORMAL OPERATION EXTERIOR CRECK

.......................

2. 1

..............

2-5

BEFORE STARTING EXGIXES

.................................. 2-1OC PRE-TAKEOFF ........................... 2-1OC =OFF .............................,. 2-14Q

TAXIING

_C_

CATEGORY "A" TAgEOFF

CLIMB

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

.............-2-1.................................. 2-14R .................................. 2-14R

CATEGORY "B" TAKEOFF

CRmSE

IN-FI;IGfff POWER ASSURANCE

APPROACH AND LANDING

...........2-14R

................ 2-14T

CATEGORY "A" APPROACH AND

.....-.................... 2-15 ............................ 2-15 GeAROUND ....................,,...,. 2-16 LANDING CATEGEORY "3" APPROACH AND LANDING

.............

AIRCRAFT WITHOUT EAPS 2-16A AnZCRAFT WIT3 EAPS .'2 .16A FAA APPROVED NOVEMBER 21. 1978

................

Revised JUNE 28. 1984

Pan 1 Table of Contents

FUGHT MANUAL

TABLE OF CONTENTS (Conr) SECTION

PAGE

POST SHUTDOWN ....................................................................... >@Minute . . Power < s A ................................. Dailv Trend Check Calcuia~ion................................................

EXTERNAL CARGO HOOK OPERATION ................................ PREFLIGHT .......................................................................... TAKEOFF .............................................................................. CARGO PICKUP ................................................................... CRUISE .................................................................................. APPROACH AND LANDING ............................................... R.H. SLIDING CABIN DOOR (MODIFIED BY KIT PIN 76070-2001 5-01 1) IKFLIGKT OPERATIOS ................... ...

LTILTPI HOIST OPERATION ..................................................... EXTERIOR INSPECTION ..................................................... 1WERlQR CHECK ................................................................ PREFLIGHT .......................................................................... YTILITY HOIST PICKUP ...................................................... HOIST DUTY CYCLE ............................................................ OPTlONAL PROCEDURES ENGINE FiRE EXTINGUISHER SYSTEM TEST

........................

FLOTATION SYSTEM TEST ........................................................

I

.

FUEL PRIMING PROCEDURES . DURING NORMAL START .....................................................

Starter mororin~...................................................................... Eneine cross prime .................................................................. Manual priming....................................................................... NAV TEST SWITCH FUNCTION ................................................. FUNCTIONAL CHECK .SNOW PROTECTION KIT P/N 7608O-3OOO8-OI2 OR 76076-30006-012..............................

FAA APPROVED NOVEMBER 2 1. 1978 Revised March 2.1989

Pan 1 Table of Contents

FUGHT MANUAL

TABLE OF CONTENTS (Cont) SECTION

PAGE

FUNCTIONAL CHECK .PITCH BIAS ACTUATOR

...................

2-21

111 EMERGENCY PROCEDURES ..................................................... 3-1

Land immediateiv ................................................................... 3-1 Land as soon as possible ...................................................... 3-1 . . ........................... 3-1 Land as soon as practicable ..................... ENGINE SYXEM MALFUNCTIONS ........................................

3-1

SINGLE-ENGINE POWER FAILURE .......................................... 3-1 SINGLE-ENGINE POWER FAILURE WITH EAPS OPERATING OR ANTI-ICE OPERATING ....................... 3-2 SINGLE-ENGINE FAILURE .HOVER (5 TO 15 FEET) ...... 3-2 SINGLE-ENGINE FAILURE ON TAKEOFF . CATEGORY "A" ...............................................................

3-2

Hover ............................................................................... -

3-2A

........................

3-2A

Prior to initial Rotation Speed of 35 KI AS

After initial Rotation During Ciimb Up to and Including the CDP (35 KlAS and 40 Feet Above Ground Allimeter Readine ( 5 5 Feer Radar Altimeter) ........................ 3-2A After CDP ........................................................................

3-2A

SINGLE-ENGINE FAILURE ON TAKEOFF . CATEGORY "B" ............................................................... 3-2B SINGLE-ENGINE FAILURE DURING CRUISE .................. 3-2B Engine Restart in Right .................................................... 3-2C Fuel Crossfeed After Engine Failure ................................. 3-2C Enpine Shutdown in Flight ............................................... 3-2D

FAA APPROVED NOVEMBER 2 1. 1978 Revised March 2. 1989

Part 1 Table of Contents

&xmR~m~i& FLIGHT MANUAL TABLE OF CONTENTS (Cont)

SECTION

PAGE

Prior to or at LDP - Balked Landing Go-Around

....-..--.-....3-2D

SINGLE-ENGINE LANDING - CATEGORY "B" ................

3-3

TAKEOFF AT 15 FEET OR BELOW .................................. 3-3 DUAL-ENGINE FAILURE DURING TAKEOFF AND J N T l A L CLIMB ................................................................. 3-4 DUAL-ENGINE FAILURE DURING CRUISE .....-............... 3-4 AUTOROTATIVE LANDINGS ......................................... 3-4A

FAA APPROVED NOVEMBER 21,1978 Revised March 2,1989

Part 1 Table of Contents

TABLE OF CONTENTS (Cont) SECTION

PAGE E N G N CONTROL MALFUNCTIONS . . . . . .

3-5

N2 CONTROL ACTUATOR RUNAWAY . . .

3-5

COUECTTVE BIAS ACTUATOR OR DROOP COMPENSATOR AMPLEER (DCA) FAILURE . . . . . . . . . . . .. .. . . . . . .

3-6

ENGINE CONTROL FAlLS TO HIGH 3-6

POWER...............................

ENGINE AUXILLARY SYSTEMS FAILURE ..

3-6A

ENGINE OIL SYSTEM FAILURE . ... . . . . . FUEL PRESSURE WARNING LIGHT ON . . FUEL FLOWMETER BEYOND LlMlTS (IF FUEL FLOWMEERS ARE INSTALLED) .. . . .. . . . . . . . . . .. .. . . . . .

3-6A 3-6B 3-6B

EAPS CAUTION LIGHT ON ..... .. .. ... . ANTI-ICE CAUTION LIGHT ON .. . .. .. .. ENGINE CHIP DEI'ECTOR CAUTION LIGHTON ..........................

MAIN GEAR BOX SYSTEM FAILURES

.....

3-7

h4AJN GEAR BOX OIL SYSTEM

FAILURES

3-7

. . . . . . - . . . . . . . . . . . . . U R E S U R E S U R E S U R E

Main Gear Box Oil Pump Failure .. . . . . .. LOW Oil RSSIE~. .. . . . . . . . . . . . . . . . . . . Hi@ Oil Te.. ........ .... .... Main Gear Box Chip Detector ...........

3-7 3-8

INTERMEDIATEOR TAIL GEAR BOX CHtP HIGH OIL TEMPERATURE DETECTED ...

28A

ROTOR BRAKE CAUTION LIGHT ON ROTOR TURNNG . . . ... . . . . . . . . . . . . . . .

3-9

FLIGHT'CONTROL HYDRAULIC SERVO SYSTEM FAILURE . . . . ..- . . . . .. . . . . . . . *

PUMP FAILURE OR LOSS OF PRESSURE IN BASIC HYDRAULIC SYSTEM . . . ... ..

29

SERVO UNIT JAM OR MALFUNCTIONING SERVO SHUTOFF VALVE . .. . ... ... ... .

3-10

FAA AF'PROVED NOVEMBER 21, 1978 Revised March 16, 1994

Part 1 Table of Contents N G H f MANUAL

TABLE OF CONTENTS (Cont) SECTTON

PAGE

PITCH BIAS ACTUATOR RUNAWAY- -

.......

3-1QA

- . ....... .

SINGLE-ACTGATOR m V E R .. . - .. 3-IOB SINGLE-ACTUATOR FAILURE.. ...... . . . 3-10B OSQLLATORY MALFUKCTIONS . 3-10B

. . -. .

OUTER LOOP SEUTDOWN ... . - . ... 3-10C 3-11 SlXGLE AXIS OUTER LOOP RAIUX)VER..

TAIL ROTOR MALFUNCTIONS .....- .- ..... 3-11 TAIL ROTOR D m SYSTEM FAILURE.. IN FORWARD FLIGHT.........-~..-- --...... . . -~

~

TAIL ROTOR D R TS~~ T E M FAILURE IN A H O n X . . .- ....- .- -... .. .... .... TAIL ROTOR CONTRCL SYSTEM FAILURE ..... .............. .- .- - - - - -

.

-

.

Phed Pitch Setting Geneml ........ Fixed Pitch SeReduced Power - - - --- . -Fixed Pitch At Hi& Power SeI .- .- -Fixed Pitch set& at cruise Power . .

--

-

-- .

- ,.....3-l2A

TAIL ROTOR CABLE FAILURE CZWIXFUNG SPFXNGINSTALL;ED

(-2'). ...........................;. . .3-1% -

FlXED PITCH m G TAlL ROTOR (=ENTERING SPRING

,

k.LfCE

E&e

................................ Comoartment Fire ...-...- .- A

- - - - - - ~ . - - -

Baggage ~ o m h e nSmoke t Detected . Smoke and Fume Elimination.. ..... ... Electrical Fire ....... ... . . . ....... ..

.

FAA APPROVED NOVEMBER 21, 1918 Revised JUNE 29, 1984

Part 1 Table of Contents FLIGHT MANUAL

TABLE OF CONTENTS (Cont) SE mori

PAGE

..- - . ..-

LAhDING GEAR MALFUNCTIONS .. ...

3-16

-

UNSAFE IhIDiCATIOh' GEAR EXTETr'SIO................................... ... 3-16 U ~ F INDICATION E GEAR RETRACTIOX . . . - . . . . . . . . . . . . . . . . - - - - - - . . . - -3-16A ---..

ELE Cl?RICAL SYSTEM FAILURES

.............

3- 16A

MASTER WARNNG AND CAUTIONADVISORY 'FAXLURE -, . . BATTERY OVERTEMPERATUEE ... SINGLE DC GENERATOR FAILURE . .DUALDC GENERATOR FAILURE.. .. DC GENERATOR OVERTEMPERATURE ( o p t i o d configaration) INVERTER FAILURE (AC GENERATOR

...... . .....

-

m

--

- - - - - - ---

-

ml\

m S R I E L D ROT CAUTTm LIGHT'-......-.-..

....... m G E D DOORS ..................... Door Jettison Procedure ............. FL H . SIDING DOOR (OPTIONAL). .... Wmdow Jettison Procedure .... ......

EMERGENCY ENTRANCES AND EXITS.. -

Rieht Sliding Door Not Modified by

k;it P/N 76070-20015-011 and all Left

FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

3-19

3-20

&R@RHKW//&

Table of Contents Part 1 List of Xllustrations

FUGHT MANUAL

TABLE O F ahTEhTS (Cant) PAGE

SECTION

DIT=G

............. ..........-....

3-21

(FLOTATIONSYSTEM)

DlTCHING PROCEDURES.. SIKGLZ- ENGDii W D H G TEE WATER (FLOTATIOK GEAR I X F U """'

.. ... EXTERNAL CARGO ROOK PROCEDURES.. ..... AFIY)FtOTATlVE LANDING TO THE WATER DITCHING (WITEOL? FLX>TATION SYSTEM

UTILITFHOISTPElOCEDURES

IV

................

3-23 3-23 3-24 3-24

PERFORMANCE DATA

........................ 4-6A EAPS EFFECT ON PERFORMANCE., .......... 4-7 ToX2uE AVAILABLE

ROTORCRAFT-EXTERNALLOAD CATEGORY 9''OPERATIONS

...............

FORWARD CLTMB PERFORMANCE

47

............

4- 13

PART I

LIST O F ILLUSTRATIONS

ITTLE 1- 1 1-2 1-3 1-4

PAGE

Category "A" b5axhxn.n Takegff Gross WeiBt Category 3'' NLaximm Takeoff

......................... GmssWeig3L. ....................... Center of Gravie Limits. ...............

Limiting Heights and Corresponding Speeds for Safe Landing Aftez an Engine Suddenly Becomes Inoperative

...

FAA APPEZOVED NOVEMBER 2 1, 1978 Revised July 26, 1982

1-2 1-3 1-4 1-6

FUGHT MANUAL

PAGE

Maximum Airspeed vs External Cargo Hook Weight Engine o r Drive System Operating Limits......... Instrument Range hlarkngs.. Vne Placards..

................ .................. ........ .....................

Installed P m e r Assarance Check Chart 70% Torque EAPS Not Insialled or EAPS InstaLled and Switch OFF InskUed Power Assurance Check Chart 75%Torque EAPS Not Insralled or EAPS InstaLled and Switch OFF Ins&lled Power Assarance Check Chart 8% Torque EAPS Not installed or EAPS Installed and Switch OFF Installed Power Assarance Check Cbart 85%Torque EAPS Not LnstafledorEAPSInstdlledand

-

-

-

...................... ......................

I-8B 1-9 1-14 1-17

2-12

2-13

-

...................... Switch OFF ......................

2-14

2-14A

Installed Power Asscvance Check Chart - 9W?& Torgae - EAPS Not

InstalledorEAPS~edand S w m i

om

......................

2-14B

bstalled Power Assurance Check Chart 95%Torque EAPS Not lnstalledorEAPSlnstalledand

-

Switch OFF

-

...................,, .

2-14C

Installed Power Asmxance Check Chart lm T o m e EAPS Not Installed or EAPS Installed and

-

-

...................... Switch OFF ......................

Switcb OFF

LPsealled Power Assarance Check W 105%TOEAPS Not Installed or EAPS Installed aad

Deleted h s & l l e d Power Asscrrance Check Chart-7O%EAPSIasWkdand SwitrhON**..*.*-....*.......-... Installed Power AssmanCe Check Chart 75%EAPS Instaled and

-

....................... - ......................*.

,Switch ON

LnstaZled Power Assurance Check C h a t 80% EAPS hstalled and Switcb ON

FAA APPROVED NOVEMBER 21, 1978 Revised JANUARY 17, 1983

2-14D

2-14E

2-14G 2-14.H 2-14J

FLIGHT MANUAL

LIST OF ILLUSTRATIONS (Cont) RGURE NUMBER 2- 13 2- 14 2- 15 2-16 2-17

2-18 2- 19 2-20 2-21 2-22 3- 1-0 3- 1 3-2 4- 1 4-2 4-3 4-4

4- 5 4-5A 4-6

TITLE Installed Power Assarance Check Chart 85% U P S Installed and Switch Oh-....,.................... Installed Power Assarance Check Chart 90%EAPS Instilled and Switch ON Installed Power Assurance Check Chart 95% UPS Installed and Switch ON Installed Power Assurance Check Chart 100%EAPS Installed and Switch ON Installed Power Assurance Check Chart 105%EAPS Installed and Switch ON 74 KIAS bstdled Power Chart.. 135 K I M - h s b U d P o w e r Chart...... 74 gIAS b s i a l k d Power Chart with EAPS h & d k d and Switch ON 135 gIAS Installed Power Chart wit31 EAPS installed and Switch ON Anti-Ice On Correction Factor

PAGE

-

- ......................... ......................... ......................... - ......................... .....

-

................................ ................................ ........ AFCS Outer Loop Shutdown ........... Electrical Fire Pmcedures (AC Gene r a t o r System)..................... Emergency Ehtmmes and Exits ....... Densitp Altitade Chart.. .............. Temperatare Conversion Chart.. ...... Airspeed Qlibration - Pilot' s System.. .......................... Ahspeed Cdlibration - Copilot' s System ............................. Torque A ' -le - 2-1/2 Minate Power OEI and 30 Minute Power OEI - EAPS Not bmabd ....................... T o m e Available - 2-112 Mimate Power OEI and 30 Minate Pawer OEf - EAPS Install4 and EAPS Switch OFT ...... T o m e Available - Takeaff Pmer. Dual Engine and NoCSuise Power, Dual Engine - EAPS Not IrskUed

...........................

FAA APPROVED NOVEMBER 21, 1978

Revised JANUARY 17, 1983

WGHT MANUAL

LIST OF ILLUSTRATIONS (Cont)

PAGE

TITLE

-

Torque Available Takeoff Power Dual Engine with EAPS installed and Switch UNh-ormal Cnrise P w e r with EAPS lastdlled and Switch OFF Category "A" Takeoff Profile.. Category "A?' Rejected Takeoff Distance Category 'A" Takeoff Distance to Achieve Takeoff Safe* Speed.. Forward Climb Performance Single m e ; 2-1/2 Minute Power. Takeoff Safety Speed.. Forrvard Climb Performance Single Engine 30-Minute Pwer, Best Rate of climb Speed.. ; Fonva& Climb Performance Single Continacms Engine, I.% Power, Best Rate of Cllimh Speed.. Categorp "A" Laxiing Prafiie.. Categorp "A" Landing Distrnce Categoxy *'B" -off Distances-. Forwad Climb Performance ?tPo Engines, Takeoff P w e r . 52 IOAS.. .. F o M Climb Performance ' l k o Engines, Maximnm Conlimous P w e r , Best Rate of Climb Speed.. .. F o M climb Performance Two Engines, Nonnal Craise Power, BestRateof CllimhSpeed CXggorg *gr9 Landing Distance Altimeter -ration Pilot' s and CopilotTsSystem E w e r Oot af Ground Effect..

....................... ....... ..........................

...... ..............

-

.............

-

.. .......

-

....... .....

-

........... .......

..............-... .........

FAA APPROVED NOVEMBER 21. 1978 Revised Nwember 16, 1981

R

R

H

Part 1, Section I

~ o p m t i n g Limitations

FUGHT MANUAL SECTION I OPERATING LIMITATIONS NOTE Compliance with the limitations in this section is required by law. WEIGHT LIMITS Maximum takeoff and landing weight is 10,500 pounds (4763 kilograms). NOTE When operated at gross weights above 10,300 pounds (4672 kilograms) the helicopter must comply with Revision 14 of the Ainvonhiness Limitations section, dated May 14, 1985, or subsequent FAA approved revisions of the Airworthiness Limitations and Inspection Requirements Manuai SA 404776-2- 1. This helicopter is to be operated using the approved loading schedule. Refer to LOADING INFORMATION in Pan 2. For minimum operating weight, refer to Figure 1-3. C.4TEGORY A OPERATIONS See Figure 1- 1 for variation of allowable takeoff gross weight with altitude and temperature. CATEGORY B OPERATIONS See Figure 1-2 for variation of allowable takeoff gross weight with altitude and temperature. CG (CENTER OF GRAVITY) LIMITS See Figure 1-3 for forward and aft center of gavity limits at various g o s s weights. Lateral CG Limits: Left or right 3 112 inches (89 mm). Left or r i ~ h 4t 112 inches 7 1 14 rnm) with hoist load attached, hover only.

FAA APPROVED NOVEMBER 2 1. 1978 Revised May 15, 1992

I

Part 1 Section I Operating Limitations FLIGHT MANUAL CATEGORY "A" MAXIMUM TAKEOFF GROSS WEIGHT ANTI-ICE OFF

GENERATOR LOAD

200 AMPS

NO BLEED-AIR

REDUCE MAXIMUM TAKEOFF GROSS WEIGHT DETERMINED FROM CHART BY THE AMOUNT SHOWN I N FOLLOWING TABLE AS APPLICABLE. WlTH COMPOUND CONFIGURATIONS. THE WEIGHT REDUCTIONS ARE CUMULATIVE CONFIGURATION EAPS INSTALLED ANTI-ICE ON ANTI-ICE ON WITH EAPS INSTALLED UTILITY HOIST INSTALLED R H SLIDING DOOR OPEN

WEIGHT REDUCTION 135 POUNDS 225 POUNDS 350 POUNDS 125 POUNDS 350 POUNDS

-MAXIMUM GROSS WEIGHT

7

6

8

li

9

GROSS WEIGHT-1000 LBS

3.'0

3-5 GROSS WEIGHT-

4.0

4:5

1000 KILOGRAMS

Fi=we 1-1 FAA APPROVED NOVEMBER 21, 1978 Revised MAY 14. 1985

S 53336 (C33)

CATEGORY "B!' M A X I M U M TAKEOFF ANTI-ICE OFF

GROSS WEIGHT

GENERATOR LOAD-200 AMPS

NO BLEED-AIR

REDUCE MAXIMUM TAKEOFF GROSS WEIGHT DETERMINED FROM CHART BY THE AMOUNT SHOWN IN THE FOLLOWING TABLE AS APPLICABLE: CONFIGURATION EAPS INSTALLED ' ANTI-ICE ON ANTI-ICE ON WITH EAPS INSTALLED

WEIGHT REDUCTION 100 POUNDS 200 POUNDS 300 POUNDS

*!LAPS SWITCH I N THE ON POSlTlON. NO WElGKF REDUCTION

WITH EAPS SWITCH I N THE OFF POSmON

GROSS WEIGHT-1000

LB

GROSS WEIGHT- 1000 KILOGRAMS

Figure i - 2 F A A APPROVED NOVEMBER 2 1.1978

Revised APRIL 13. 1988

Part 1, Section I Operating Limitations

AIKHRSK~~~& FLIGHT MANUAL

LOADING LIMTTS Maximum allowable cabin floor loadin2 is 75 pounds per square foot (366 Kg per square meter). Tne maximum allowable floor loading for the bagsage compartment is 75 pounds per square foot (366 Kg per square meter) for weizhts up to the maximum compartment capacity of 600 pounds (272 Kg)

TYPES OF OPERATION CATEGORY A and B

I

Transport Day, Night, VFR Day?Nizht IFR.The helicopter is not considered airworthy for operation under Instrument Meteorological Conditions ( M C ) unless the following equipment is installed and operatins. 1. AFCS (minimum requirement is for one operative pitch, roll, and yaw

channel). 2. Copilot flight controls and flight instruments.

FAA APPROVED NOVEMBER 2 1: 1978 Revised June 17.1997

FUGHT MANUAL

Part 1 Section I Operating Limihtions

FAA APPROVED NOVEMBER 21, 1978 Revised M a y 14, 1985

FLIGHT MANUAL

Part 1, Section I Operating Limitations

AC Power - AC Generator and inverter. Dual 5" Vertical Gyro Indicators or equivalent. C- 14A Com System

Deleted Cyclic Stick Trim An operative navigation and communication system that has demonstrated compliance with the pertinent airworthiness regulations and also meets the requirements of the applicable operating regulations.

I

Two DC Generators. Overwater operations provided that the emergency flotation gear, PM 76-076-02002, and suitable lifesaving equipment (life jackets, rafts, etc.) as required by the operating rules and FAR 29.141 1,29.1415, and 29.1 1561 are installed. CATEGORY "BnROTORCRAFT- EXTERNAL LOAD COMBINATIONS Class "B" external loads NOTE A class "B" external load is an external load that can be jettisoned and is lifted fkee of land or water during rotor-craft operation. MINIMUM FI IGHT CREW

Visual Meteorological Conditions - 1 pilot (Single pilot operations not permitted from left seat). NOTE Aircraft Serial No. 760001 and 760002 require installation of E.O. 72859 against 76550-02001 and E.O. 71 195 against 76080-45010 to be eligible for 1 pilot operation. Lnstrument Meteorological Conditions - 2 pilots.

FAA APPROVED NOVEMBER 2 1,1978 Revised November 6,2003

Part 1, Section I Operating Limitations

A !FLIGHT ~ ~ MANUAL Z Z ? B H ~ ~ ~ ~

LIMITING HEIGHTS AND CORRESPONDING SPEEDS FOR SAFE LANDlNG AFTER AN ENGINE SUDDENLY BECOMES INOPERATIVE THESE CURVES ARE APPLICABLE TO ALL ALTlTUDES AND TEMPERATURES AT THE CORRESPONDING MAXlMUM ALLOWABLE TAKE OFF GROSS WEIGHT AS DETERMINED FROM FIGURES 1 - 1 AND 1.2. THE HIGH HOVER POINT IS BASED ON MAXIMUM OGE HOVER WEIGHT AND HAS BEEN DEMONSTRATED AT 10.500 POUNDS.

I

INDICATED AIRSPEED

HNOTS

ABOVE 10,300 LBS'

1.56398 tC33)

Figure 1-4

FAA APPROVED NOVEMBER 2 1,1978 Revised May 14,1985

AIRSPEED 3.ams

NOTE The S-76A nses an Aero Mechanisms Part No. 8502CS20LW, Aerosonic Part No. 20020-11190 or Aerosonic Part No. 20020-11293 airspeed inditator.

See , V placard, Figure 1-7, for nrhfionof ,V tempeatate, and pressare al-e.

wfth gmss we=

FAA APPROVED NOVEMBER 12, I978 Revised March 17, 1986

,,V power off is 141 KIAS. See Vne placard, Figure 1-7, for variation of Vne with temperature and pressure altitude. Maximum airspeed during OEI operations with N, below 100%is the best rate of h b speed. With usable fuel per tank indicating 80 pounds or less, maximum allowable airspeed is 1 2 6 KIAS. Maximum airspeed with landing gear down or in transit is 130 KIAS. Maximum airspeed for windshield wiper operation is 141 KIAS. Maximum groundspeed for landing, takeoff, or taxi is 40 knots. M m u m groundspeed for brake application is 34 Imots. ALTITUDE LIMITS Takeoff and landing:

6900 feet density altitude.

11,000 feet density altitude for helicopters modified by kit P/N 76070- 30005 t o relocate engine inlet temperature sensor. Enroute:

15,000 feet density altitude.

AMBIENT TEMPERATURE LIMITS -34.4OC (-30°F) to. ISA + 36.7OC not to exceed 48.g°C (120°F).

FLIGHT LIMITS See Figure 1-4 for altitude and airspeeds to be avoided at low altitude in case of engine failure.

No aembatic maneuvers allowed. 360° hovering turns in less than 1 2 seconds prohibited. flight in known icing conditions prohibited.

FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

1

Part 1 Section I &!~~~BBsKw~& Operating Limitations ~ G MAIJUAL M Maximum airspeed for sideward flight or crosswind hover is 35 knots, up to and including 6900 feet density altitude. Maximum airspeed for sideward flight or crosswind hover is 1 7 knots, from 6900 feet up to and including 11,000 feet density altitude. Maximum airspeed for rearward flight or tailwind hover is 35 knots, up to and including 6900 feet density altitude. Maximum airspeed for rearward flight or tailwind hover is 1 7 knots, from 6900 feet up to and including 11,000 feet density altitude.

Do not re-engage a particular AFCS channel after a known malfunction exists in that particular channel. Flight in falling. o r blowing snow is prohibited unless aircraft is fitted k t h snow protection kit P I N 76080-30008 o r P I N 76076-30006-012 and a satisfactory functional check has been performed. Flight in falling and blowing snow with EAPS installed is prohibited. Cockpit ventilation must be provided by any of the following: 1.

Overhead vents

2.

Pilot's window

3.

Heater blower

4.

Heater bleed-air

5.

ECU

Fuel crossfeed operations limited to: Fuel crossfeed operation is prohibited in flight except in strict accordance with the emergency procedure, Fuel Crossfeed After Engine Failure on page 3-2C'. :

Fuel crossfeed operations with vent 3ine check valves (Xt PIN 76070- 3QO2SbOll)installed limited to: Category "Am-emergency operation only

1-8

Category "Bn-cruising flight only FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

[email protected] FLIGHT MANUAL

Part 1, Section I Operating Limitations

Minimum IMC airspeed is 60 KIAS (AFCS Phase 11). .Minimum IMC airspeed is 50 KIAS (AFCS Phase 111). Steepest Demonstrated Approach Gradients (AFCS Phase 111). 7.5' at 60 to 80 KIAS 6.5' at 60 to 125 KIXS Ensine ANTI-ICE ON at 2°C and below, with visible moisture. External door locks must be unlocked before flizht. SLIDING CABIK DOOR LIMITS Right sliding cabin door. unless modified by Kit PIN 76070-20015-011 must be closed for all flight operations.

1

Left sliding cabin dodr and right sliding cabin door modified by Kit PIN 7607020015-01 1 may be opened for flight operations with these restrictions:

1

Maximum airspeed for opening and closing slidins cabin door in flight is 5 0 K1.G and is limited to level f l i ~ h or t descent. Maximum airspeed with sliding cabin door in full open and locked position is 74 KIAS. IMC flight prohibited with sliding cabin door open. An operable flashlight must be available in the cabin if the sliding door is to be opened in flight at night. FLOTATION SYSTEM LIMITS Maximum airspeed for inflation of the emergency flotation e a r is 75 KIAS. Maximum airspeed with emergency flotation gear inflated is 75 M A S . Maximum water speed with emergency flotation system inflated is 33 knots. Landing :ear must be down prior to float inflation.

FAA APPROVED NOVEMBER 2 1, 1978 Revised May 15, 1992

Part 1 Section I Operating Limit ation

~I/K~D?S/KOY~~~ FLIGHT MANUAL

Maximum demonstrated airspeed for sideward flight or crosswind hover is 20 knots. Maximum pressure altitude with emergency flotation gear inflated is 5000 feet. EXTERNAL CARGO HOOK LIMITS External hook operations must be conducted by a qualified flight crew under the provisions of the operating rules for rotorcraft external load operations for loads that are jettisonable and are lifted free of the surface (Class B loads). Normal operations are permitted with the external cargo hook installed, but not used. The height-velocity diagram in Figure 1-4 does not constitute a limitation when conducting rotorcraft-external load operations. WEIGHT LIMITS Maximum weight of the external cargo is 3300 pounds (1500 k g ) . AIRSPEED LIMITS

Maximum airspeed with an external cargo weight of 1900 pounds (863.6 kg) or less is 136 KIAS. Maximum airspeed must be reduced with increases in external cargo weight. Refer to Figure 1-4A. NOTE It should be noted that operational use of the cargo sling is a highly specialized industrial procedure. Extreme caution must be taken to be sure that the loads carried and the speed range throughout which operation is intended do not adversely effect the controllability characteristics of the helicopter. It is the responsibility of the operator to establish safe and sound limitations for each operation.

FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

Part 1, Section I Operating Limitations

FLIGHT MANUAL NOTE The external carso hook is located forward of the main rotor shaft. Releasing a load will cause the aircraft to pitch nose UP-

LTILITY HOIST LIMITS Utilie hoisr operations must be conducted by a qualified flight crew under the provisions of the operating rules for rotorcraft external load operations for loads that are jettisonable and are lifted free of the surface (Class B loads). Normal opelatiom are permitted with the utility hoist installed, but not used.

FAA APPROVED NOVEMBER 21. 1978 Revised March 19, 1996

Part 1. Section I ~ I / x ( ~ ~ R H K OOperatins ~& Limitations FLIGHT MANUAL

MAXIMUM AIRSPEED V S DENSITY ALTITUDE AND EXTERNAL CARGO HOOK WEIGHT

0

20

40

60

AIRSPEED

80

100

- KlAS

Figure 1-4A

FAA APPROVED NOVEMBER 21, 1978

Revised March 19, 1996

120

1 40

Pan 1. Section I Ooeratine Limitations w

FLIGHT MANUAL

NOTE It should be noted that operational use of the utility hoist is a highly specialized procedure. Extreme caution must be taken to be sure that the loads camed and the speed range throughout which operation is intended do not adversely effect the controllability characteristics of the helicopter. It is the responsibility of the operator to establish safe limitations for each operation. The height-velocity diagram, Figure 1-4. does not constitute a limitation when conducting rotorcraft-utility hoist operations. A cabin intercom system with operational "hot mike" capability is required for hoist

operations in which loads are brought into or out of the cabin. The last 20 feet of hoist cable (colored red) is unusable. WEIGHT LKMITS The maximum utility hoist load is 600 pounds (272 kg). The maximum weight of the rotorcraft - external load combination is 10,500 pounds (4763 kg). Maximum baggage compartment load is limited to 600 pounds (272 kg). The weight to the left (port) or right (starboard) of the centerline of the baggage compartment should not be over 300 pounds (!36 kg). The hoist load may be restricted by lateral loading limits. See hoist loading information, Part 2, Section 11. lMINIMUM FLIGHT CREW Minimum flight crew for utility hoist operations is: 1. Two pilots, and a hoist operator in communication with the pilots.

I

2. One pilot and a hoist operator in communication with the pilot if the hoist configuration includes a collective mounted hoist upidown switch in addition to the cabin mounted provisions. AIRSPEED LIMITS With a load on the utility hoist, the maximum airspeed for sideward flight or crosswind is 30 knots.

1-8D

FAA APPROVED NOVEMBER 21, 1978 Revised May 15, 1998

Part 1 Section I A///K@RH~~& Operating Limitations FUGHT MANUAL

With ioad on the utility hoist, the maximum airspeed in forward flight may be restricted by stability of the load and is not to exceed 74 KIAS. ENGIKE LIMITS (See Figure I- 5 ) FUEL Primary Fuel ASTM D -1655 Jet A, A l , or B MIL-T-5624 Grade JP-4 or JP-5 MIL-T-83133 Grade JP-8 Cold Weather Fuel (Below 4°C (40°F)) MIL-T-5624 Grade JP-4 *MIL-T-5624 Grade JP-5 with anti-ice additive conforming to MIL-I27686 (Philips Petroleum Company MB-55 or Equivalent) in concentration of 0.035 to 0.15% by volume. ASTM D-1655 Jet B *ASTM D-1655 Jet A, or A1 with anti-ice additive conforming to MIL-I27686 (Philips Petroleum Company MB-55 or Equivalent) in concentration of 0.035 to 0.15% by volume. **AVGAS/jet fuel mixture (alternate fuel)

FAA APPROVED NOVEMBER 21,1978 Revised September 7, 1990

Part 1 Section I Operating Limit ations

&~K
NOTE *Starting may not be consistent below 4OC (40°F). **Mixture Ratio: 1 Part AVGAS, Grade 8018'7, to 2 Parts Jet Fuel (Jet A , Jet Al, or JP-5) by volume may be used for unrestricted periods of time. AVGAS, Grade 100/130 (100L) with a maximum of 2.0 ml/Gal. ( 2 . 4 rnU1mperia.l Gallon, -528 mllliter) lead content may be used in place of grade 80187 in the same proportions with jet fuel for not over 300 hours during any overhaul period. Do not use above 4OC (40°F). Do not use AVGAS containing Tricresyl-Phosphate (T). If the jet fuel1AVGAS mixture is added to JP-4 or Jet B y add. anti-ice additive in concentration of 0 -035 to 0.15% based only on the jet fuel1AVGAS volume added. If the jet fuel to be mixed with AVGAS is JP-5, Jet A, or Jet A 1 to which anti-ice additive has not been added; add anti-ice additive in concentration of 0.035 to 0.15% based on entire fuel volume.

FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

flw 1 k c t ~ o I~ ) ( > p c r ; ~ t n gL i ~ ~ ~ ~ ~ ~ i l l o l - , ~ FLIGHT MANUAL

Part 1 Section I Operating Limitations

&!f%B~~!ifl& NGHT MANUAL

-

MIL-L-23699 or later revision Low temperature limit for stark: of -40°C (-40°F) MIL-L-7808For later revision - For starting temperatures between -40" w -54OC (-40' to -65°F) NOTE If sny of b e following limits a r e exceeded, refer to ' e n a e manufacmer's maintenance manual for action to be a e n .

OIL TEMPERATGRE IJMrrs

Maximum: 107°C (225"F) Minimum flight 0°C (32°F) (provided that engine oil pressure is within specified limits)

Maximum: 130 PSI Minimum: 50 PSI

FlSEL PRESSURE LIMITS (IF F'UEL PRESSURE GAGES ARE IIU'STALLED)

Minimum: 5 PSI

mEL FLOW LlMITs (IF FVEL FLOWMETERS ARE INSTALLED) Minimum: 75 pounds per hour

-urn:

412 pomds per hour NOTE Fuel flow may go beyond these limits during transient operanons including start.

E N G W TORQUE IJMITS

XOTE Refer to TRANSMISSION TORQVE L I h C E in this section for additional torque limits.

104.6% T o q e

- 5-minute takeoff and maximum continuous limit, dual-engine (Refer to Engine Ratings and Recommended Usage, Section I Part 2 for recommended usage of maximum continuous torque. )

F A 4 APPROVED NOVEMBER 21, 1978 Revised JA-.?TARY 17, 1979

- 30-minute limit, single-engine 111.2%.~ o r q a e- 2- 1/2 minute limit, single-engine 111.2% to 155% Torque - 16-second transient, s@e 1 0 4 6% Torque

engine

N1 (GAS PRODUCER) SPEED =TS

- Maximum for continuous operation 105% to 106%N1 - 10 second transient 105% N1

N2 (POWER TURBINE) SPEED LIMITS 114% N2

- Maximum at 0%

torque varying linearly to 107.1% 32 a t

111.2% torque

Transient: 15 second limit varies Linearly from 119% at flight autorotation to 109%at 111.2% tmque.

FAA APPROVED NOVEMBER 21, 1978 Revised FEZ3RUARY 1, 1979

Part 1 Section I FUGHT MANUAL

Tj (POWER TURBINE INLETTEMPERATURE) LIMITS Steadv State Limits 826°C - 2- 1/2-minute power. one engine inoperative 798°C - 30-minute power. one engine inoperative 768°C - Eminute takeoff and maximum con:inuous (Refer to Engine R a t i n s and Recommended bsage Section 1, Part 2, for recommended usage of maximum continuous Ts. Starting znd Shutdown 826" to 927°C - Starting transient 10-second limit, time at 927°C is limited to 1 second. Power Transients 826" to 905°C - 12 seconds

To prolong engine life. avoid this transient condition ir! normal operations. TR.4NSMISSIOh LIMITS TORQUE LIMITS (See Figure 1-5.) Dual-EnGne Operation 1000~;Torque per engine - 5-minute takeoff and maximum continuous.

Takeoff torque may exceed 100% on one engine to a maximum of 104O/11,provided that the torque on the other engine is less rnan 96O/0 and rhe sum of the individual torque values does not exceed 200%:. Single-En$ne Operation 100°/oTorque - Maximum continuous 104OA1Torque - 30-minute limit FAA APPROVED KOVEMBER 21,1978 Revised September 7, 1990

Part 1, Section I Operating Limitations

&#K!@RS~~E F U G M MANUAL

115% Torque - 2 112 minute limit 11 5 8 to 155% Torque - 16 second m s i e n t

MIL-L-21260, Type I, Grade 30 - Low temperarure limit -9°C (15°F)

Dexron II or E l ATF - Low temperarure Iirnit -34°C(-30°F) MIL-L-23699 - Low temperarure limit -34°C (-30°F)

OIL TEMPERArn L m s

Minimum: -20°C OIL PRESSURE LIMITS Maximum: 120 PSI Minimum: 20 PSI ROTOR LJMITS POWER OFF Maximum: 115% Nr

Minimum: 87% Yr

POWER ON Maximum: 107% Nr Minimum: 100% Nr - (dualen-&e operation) Minimum: 96% 'cr

- (one engine inoperative)

96% to 99% Nr limited to one en-&e operation up to best rate of climb speed or dual-en,oine operation up to Vne when required for execution of the appropriate emersency procedure. Transient: 82% Transient: 70% Nr, at touchdown during single-ensine landings and rejected takeoffs.

1-12

FAA APPROVED XOVEMBER 21, 1978 Revised September 12, 1995

A H K D R . K O ~ ~ Part ~ 1 Section I FLIGHT MANUAL

ROTOR BRAKE LIMITS ROTOR STOPPED Engine operation (one or two) limited to idle with rotor brake on.

ROTOR TURNING Rotor brake application limited to one engine only operating at idle or both engines shut down. Maximum rotor speed for normal rotor brake application is 65% Nr. Maximum rotor speed for emergency rotor brake application with both engines shut down is 107% Nr. A rotor shutdown using the rotor brake shall not be performed more than one time in any 10-minute period.

ROTOR BLADE LIMITS - ROTOR STOPPED DUAL ENGINE OPERATION Main rotor blades must be positioned at approximately 45' to the longitudinal axis of the helicopter.

SIXGLE-ENGINE OPERATIOS At ambient temperatures of 37°C (100°F) and below with no tailwind component, unlimited at any rotor position. At ambient temperatures above 37°C (lOO°F), the main rotor blades must be positioned at approximately 4 j 0 to the longitudinal axis of the helicopter.

ENGINE COWLING REMOVED OR HINGED BACK Single-engine operations permitted only with operational rotor brake applied with main rotor blades positioned a t 4 j 0 to the longitudinal axis of the helicopter and nose of the helicopter positioned into the wind.

See Figure 1-6.

FAA APPROVED NOVEMBER 21,1978 Revised March 2. 1989

Part 1 Section I Operating Limitations

&FLIGHT RB? B?SB~E MANUAL

PLACARDS "THIS HELICOPTER MUST BE OPERATED IN ACCORDANCE WITH THE OPERATING LIMITS SPECIFIED IN THE FAA APPROVED ROTORCRAFT FLT MAN. THE AIRWORTHINESS LIMITATIONS SECT OF THE ROTORCRAFT MAINTENANCE MANUAL MUST BE COMPLIED WITH." - on instrument "CAUTION - TURN OFF STROBE LIGHT WHEN TAXIING NEAR OTHER AIRCRAFT OR WHEN FLYING IN FOG OR CLOUDS. STANDARD POSITION LIGHTS MUST BE USED FOR ALL NIGHT OPERATIONS." - on instrument "CAUTION - BELOW 80 POUNDS FUEL REMAINING PER TANK, REDUCE AIS TO 126 KIAS OR LESS." - on instrument Vne PLACARDS see Figure 1-7. - on instrument

CAUTION ENGINE OPERATION WITH ROTOR STOPPED LIMITED TO: BLADES 45" TO LONG AXIS

OTHER BLADE POSITIONS

UNRESTRICTED

NO TAIL WIND COMPONEhT AND ONLY ONE ENGIEiE AT IDLE AND ONLY 37°C OAT OR BELOW

- on instrument

1-12B

FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

Part 1 Section I ABK@BH~Y~&Operating Limitations FUGHT MANUAL

"THE MAXIMUM ALLOWABLE FLOOR LOADING FOR THE BAGGAGE COMPT. IS 75 LB. PER SQ. FT.FOR WEIGHTS UP TO THE MAXIMUM COMPARTMENT CAPACITY OF 600 LB." - on baggage companment doors

I

CAUTION "NORMAL ROTOR BRAKE STOPS ARE AUTHORIZED ONLY AT 65% Nr OR LESS WITH BOTH ENGINES SHUT OFF, OR ONE ENGINE AT IDLE." - on rotor brake handle of helicopters equipped with manual rotor brake "CARGO HOOWHOIST LIMITS ROTORCRAFT-LOAD COMB. CLASS B NO PERSON MAY BE CARRIED UNLESS: H E IS A FLIGHT CREWMEMBER. H E IS A FLIGHT CREWMEMBER TRAINEE. HE PERFORMS AN ESSENTIAL FUNCTION IN CONNECTION WITH THE EXTERNAL LOAD OPER. HE IS NECESSARY TO THE ACCOMPLISHMENT OF THE WORK ACTIVITY DIRECTLY ASSOCIATED WITH THAT OPER." - on instrument "CARGO HOOK - MAX. CAP. 3300 LB 1496.88 KG" - on bottom skin forward of cargo hook opening "PILOTS PERMISSION REQUIRED TO OPEN SLIDING DOOR. INSURE THAT ALL ENGER SEAT BELTS/HARNESSES ARE FASTENED." on 76076-20029 sliding doors or on doors modified by P/N 76070-20015-01 1 or -012

"CHECK SLIDING DOOR LOCKED OPEN BY ATTEMPTING TO PULL IT CLOSED WITHOUT PULLING ON HANDLE. INSURE PINS ARE ENGAGED." - on aft door frame of sliding doors modified by P/N 7607020015-01 1

"MAXIMUM LOAD 600 LB."

- on hoist

FAA APPROVED NOVEMBER 21,1978 Revised September 7, 1990

I I

I I I

Part 1 Section I Operating Limitations

FLIGHT MANUAL

TRIPLE TACHOMETER

Wr TACHOMETER (ROTOR) 15% MAXIMUM POWER OFF 107% TO 115% PRECAUTIONARY RANGE

11 107%. MAXIMUM PO'NER ON 1 100% TO 107% - NORMAL RANGE 1 8 7 TO 100%. PRECAUTIONARY RANGE

TORQUEMETER 11 1.2% - 2-1/2 MINUTE LIMIT. SINGLE ENGINE

87% - MINIMUM

104% TO 111.2% PRECAUTIONARY

N2 TACHOMETER (ENGINE O N L Y ) 114% - MAXIMUM 17% TO ll4'K PRECAUTIONARY RANGE 95% TO 107% NORMAL RANGE

RANGE

104% - 30 MINUTE LIMIT SINGLE ENGINE - T O 104% - PRECAGTIONARY RANGE 100%. TAKEOFF 5 MINUTE LIMIT DUAL ENGINE AND MAXIMUM CONTINUOUS LIMIT i 86.3% TO 100% - ADVISORY AREA* 0% TO 100% - NORMAL OPERATION

100%

' NOTE: THE GREEN AND WHITE ARC ON THE TORQUEMETER REPRESENTS RECOMMENDATIONS OF THE ENGINE MANUFACTURER. SEE PART 2. SECTION I FOR DETAILS

W 1 TACHOMETER 105"" MAXIMUM 59% TO 105"~,,NORMAL RANGE 59'" MINIMUM

Fik-ura 1 - 6 (Sheel 1 o l 3) r':\:\

:\PPIIOVED NOVEMBER 2 1 , 1978 Revised July 2 6 , 1982

P a r t I Section 1 Operating Limitations

FLIGHT MANUAL

ENGINE OIL TEMPERATURE/PRESSURE 1 0 7 T - MAXIMUM 0°C TO 107°C - NORMAL RANGE B O°C . MINIMUM FOR FLIGHT 130 PSI - MAXIMUM & -.. 90 TO 130 PSI. NORMAL RANGE 50 TO 90 PSI . PRECAUTIONARY RANGE 50 PSI - MINIMUM

TRANSMISSION OIL TEMPERAlURE/PRESSURE 105O TO i20°C. PRECAUTIONARY RANGE J: 15O TO 1 0 5 % . NORMAL RANGE .20° TO 15OC . PRECAUTIONARY RANGE -20°C - MIN!MUM R 120 PSI - MAXIMUM 70 TO 120 PSI . PRECAUTIONARY RANGE E 3 0 TO 70 PSI . NORMAL RANGE 2 0 TO 30 PSI . PRECAUTIONARY RANGE -3 PSI . MINIMUM

FUEL PRESSURE 5 TO 50 PSI . NORMAL RANGE 5 PSI . MINIMUM

AIRSPEED INDICATOR TO 30 KlAS - PRECAUTIONARY RANGE. INDICATIONS INCONSISTENT 30 TO 155 KlAS . NORMAL RANGE 155 KlAS - MAXIMUM

HYDRAULIC PRESSURE 3300 PSI . MAXlMlJM 2800 TO 3200 PSI . NORMAL RANGE

-

2700 PSI . MINIMUM

Figure 1-6 (Sheet 2 of 3 )

FAA A P P R O V E D NOTJEM!3ER 21, 1978 R e v i s e d J U N E 2 8 , 1984

P a r t i Section I Operating L i m i t a t i o n s

FLIGHT MANUAL

T 5 INDICATOR

- 2-1/2 MINUTE LIMIT SINGLE ENGINE 798" TO 8 2 6 ' C . PRECAUTIONARY RANGE 798'C - 3 0 MINUTE L I M I T . SINGLE ENGINE 768" TO 798'C - PRECAUTIONARY RANGE 768°C - TAKEOFF 5 MINUTE LIMIT DUAL ENGINE AND MAXIMUM CONTINUOUS LIMIT r.694" TO 768" C - ADVISORY AREA 350" TO 768'C - NORMAL OPERATION 1 826°C

LIGt!T INDICATIONS

0 0 0 0 I 0 0

a

O 0

a

-

350" AND 695'C AND O 769'CAND 0 799°C AND 827" C AND

NOTE THE GREEN AND WHITE ARC ON THE T5 INDICATOR AND GREEN AND WHITE LIGHT ON THE T5 REPEATER INDICATOR REPRESENTS RECOMMENDATIONS OF THE ENGINE MANUFACTURER. SEE PART 2 SECTION 1 FOR DETAILS.

FUEL FLOWMETER m 4 1 2 LB/HR-MAXIMUM B 75 TO 4 1 2 LB/HR . NORMAL RANGE

m 75 LB/HR

- MINIMUM

F i z ~ r e1-6 (Sheet 3 of 3) FAA APPEIOVED NOVEMBER 21: 1978 Revised J u l y 26, 1982

ABOVE ABOVE ABOVE ABOVE ABOVE

Vne PLACARDS

Vne POWER-ON. TAKEOFF GROSS WEIGHT 8750 POUNDS AND BELOW

Vne

POWER-ON.

TAKEOFF G R O WEIGKT ~ 8%1 POUNDS

tOUISOo=

Vne POWER-OFF

ALL WElGKIS

F'igure 1-7 FAA APPROVED NOVEMBER 2 1, 1978 Revised MAY 14, 1985

Part 1. Section I1 Normal Procedures

FLIGHT MANUAL SECTION II NORMAL PROCEDURES EXTERIOR CHECK The pilot will determine that the following exterior preflight check has been done before the first flight of the day or before the next flight after extended maintenance. Those items marked with an asterisk (*) should be checked just prior to each flight. Following removal and proper stowage of protective plugs, tie downs, and other associated supplementary equipment, the following check is performed beginning at the pilot's door and proceeding clockwise around the helicopter.

CHECK *Pilot and Right Cabin Doors

Cyclic Base Covers (Boots) Rotor Brake Switch Battery Switch Rotor Brake Switch Manual Rotor Brake Nose Flotation Gear Bottle *Right Cabin Sliding Door

Main Gear Box (right side)

FOR Proper operation, fit, condition, window clean and unscratched, condition of hinges and latches, security of jettison handles. Ensure keylocks unlocked (if installed) by physically opening each door from the outside Check clear of interference with controls ON ON momentarily; check that ROTOR BRAKE caution light is on. REL OFF Proper pressure Proper operation, security, condition of window and sliding rails, security of window jettison handle. On doors modified by Kit PN 76070-200 15-011, check that upper and lower door track springs are installed and operating properly. Proper oil level

FAA APPROVED NOVEMBER 2 1,1978 Revised May 25,2004

I

Part 1, Section I1 Normal Procedures

FLIGHT MANUAL

FOR

CHECK Oil Filter By Hydraulic Module - Second Stage Main Rotor Servo (if aircraft has been nonoperational for more than 2 hours in freezing temperatures)

Unpopped Proper fluid level, filter button, no leakage Attempt to manually move input link to Forward Main Rotor Servo. Input link should move freely with no restrictions approximately !4 inch. Main Rotor Head General condition Main Rotor Damper Reservoir Proper fluid level Main Rotor Blades General condition: tip cap for condition and security; blade upper and lower surfaces for raised skin indicative of disbond. Rotate the rotor system as required to view all blade surfaces and tip caps. *Engine Oil Reservoirs Proper fluid level. If low and engine has been stopped more than 15 minutes, motor engine 30 seconds and recheck. When securing filler cap, make sure latching mechanism is properly engaged before closing access . CAUTION Normal engines use a minimal amount of oil. Any sudden increase in oil consumption is indicative of oil system problems which must be corrected. Engine Oil and Fuel Filter By-es

Unpopped NOTE

It is possible for the impending oil filter by indicator to extend during a start of a cold soaked engine, giving an erroneous indication of a dirty oil filter. If the impending filter by indicator is extended, run the engine until the oil is at operating temperature and push the indicator button in. If the button remains in throughout the normal speed range of the engine, the filter does not require cleaning.

FAA APPROVED NOVEMBER 2 1,1978 Revised September 23, 1999

Part 1 Section I1 Normal Procedures

FLIGHT MANUAL

FOR -

CHECK Engine External Oil Filter Byes (optional)

Unpopped

I

Engine Fire Bottle

Proper pressure

I

Engine Area

General condition of engine. Check all accessible areas for obvious loose bolts, broken or loose connections, security of mounting accessories, broken or missing safeties, and evidence of fuel and oil leakage.

Compressor Inlet

Damage, dirt accumulation

I

NOTE Check compressor inlet carefully when helicopter.is operated in sand, dust, snow, or icing environments. Oil Cooler Blower Duct

Obstructions

Engine Air Intake

Obstructions and foreign objects which may be drawn into compressor

I

EAPS and EAPS doors

Condition, cleanness, security

I

'Fuel Filler Cap

Fuel level, cap secure

*Baggage Compartment

Contents, fuel leakage, door secure

FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

,

2-2A

Part 1, Section I1 Normal Procedures

FLIGHT MANUAL

FOR

CHECK Electrical Compartment

For hoist mission check, HOIST POWER and HOIST CONTROL circuit breakers in

*Electrical Compartment Access

secure

Main Landing Gear Compartment

Tire condition and inflation, strut extension, door and gear actuator rods properly attached and undamaged, condition of flotation bag sharp edge protection, fluid leakage, elastomeric bearing condition. Tiedown ring stowed flat against shock strut

*Tiedown ring

CAUTION Tiedown rings that are not stowed flat may lead to failure of the landing gear to retract or extend. *Downlock pin

Downlock pin removed

Flotation Gear

Bags for proper storage, covers for condition and security, bottles for proper pressure

Hydraulic Lines

Leakage

Fluid Drains and Vents

Leakage

Antennas Tail Cone Static Ports

Damage, corrosion, secured Clear of soot and obstructions

Pylon Skin and Fairings

General condition

Horizontal Stabilizer

General condition. Check that the attachment is tight and that there is no movement (play) in any direction

Side Position Light

Cracked or broken lens

Intermediate Gear Box Fairing

Lower surface for scrapes and condition, air intake for obstructions

*Intermediate Gear Box

Proper oil level

Tail Rotor Head

General condition

Tail Rotor Blades

General condition

*Tail Rotor Gear box

Proper oil level

Anti-collision and Position Light

Cracked or broken lenses

FAA APPROVED NOVEMBER 2 1,1978 Revised May 25,2004

I

Part 1, Section I1 Normal Procedures

As~m~srnflE FLIGHT MANUAL FOR -

CHECK Pylon Skin and Fairings (Left Side)

General condition

Horizontal Stabilizer

General condition

Side Position Light

Cracked or broken lens

Antennas

Damage, corrosion, secured

Tail Cone Static Ports

Clear of soot and obstructions

Tail Cone Access

Security, condition

Main Landing Gear Compartment

Tire condition and inflation, strut extension door and gear actuator rods properly attached and undamaged, condition of flotation bag sharp edge protection, fluid leakage, elastomeric bearing condition.

I

*Tiedown ring

Tiedown ring stowed flat against shock strut CAUTION

Tiedown rings that are not stowed flat may lead to failure of the landing gear to retract or extend.

*Downlock pin

Downlock pin removed

Flotation Gear

Bags for proper storage, covers for condition and security, bottles for proper pressure Leakage Leakage Drain - Check for water and visible contaminants. Adequate water drainage is provided only with the helicopter approximately level (less than 3' nose up or nose down) Contents, fuel leakage, door secure Fuel level, cap secure Unpopped General condition of engine. Check all accessible areas for obvious loose bolts, broken or loose connections, security of mounting accessories, broken or missing safeties, and evidence of fuel and/or oil leakage

Hydraulic Lines Fluid Drains and Vents Fuel Sump Drain

*Baggage Compartment *Fuel Filler Cap Thermal Relief Indicator Button Engine Area

Engine Fire Bottle

Proper pressure

Engine Oil and Fuel Filter Byes

Unpopped

FAA APPROVED NOVEMBER 2 1,1978 Revised May 25,2004

Part 1, Section II Normal Procedures FLIGHT MANUAL

FOR

CHECK NOTE

It is possible for the impending oil filter by indicator to extend during a start of a cold soaked engine, giving an erroneous indication of a dirty oil filter. If the impending filter bfiass indicator is extended, run the engine until the oil is at operating temperature and push the indicator button in. If the button remains in throughout the normal speed range of the engine, the filter does not require cleanhg

.

Compressor Inlet

Damage, dirt accumulation NOTE

Check compressor inlet carefully when helicopter is operating in sand, dust, snow, or icing environments. Engine Air Intake

Obstructions and foreign objects which may be drawn into compressor

EAPS and EAPS doors

Condition, cleanness, security

Oil Cooler Blower Duct

Obstructions

Main Rotor Head

General condition

Hydraulic Module - First Stage

Proper fluid level, filter bunon, no leakage

Main Rotor Servos (If aircraft has been nonoperational for more than 2 hours in freezing temperatures)

Attempt to manually move input links to L a W and Aft Main Rotor Servos. Input links should move freely with no restrictions approximately 114 inch.

WARNING Specifically check the security of the aft engine cowl latches and straps just prior to each fight. *All Engine and Transmission Doors and Cowls

Condition, security including hinges, latches, camlocs, and sfraps

FAA APPROVED NOVEMBER 21,1978 Revised September 23, 1999

I

Part 1, Section II Normal Procedures FLIGHT MANUAL

FOR -

CHECK Electric Rotor Brake Module

Proper fluid level, leakage

Manual Rotor Brake Accumulator

Proper fluid level, leakage

Rotor Brake Pucks (2)

Proper disc clearance

Nose Gear Lockpim (copilot's side of console)

Removed

.

*Copilot and Left Cabin Door

Proper operation, fit, condition, security, window clean and unscratched, condition of hinges and latches, security of jettison handles. Ensure keylocks unlocked (if installed) by physically opening each door from the outside.

*Left Cabin Sliding Door

Proper operation, security, condition of window and sliding rails, security of window jettison handle

*Copilot's Seat Belt

Security, if to be unoccupied

Nose Flotation Gear Bottle

Proper pressure

Windscreen

Glass clean and unscratched

Windshield Wipers

Clean, wear, condition, security

Nose Gear Compartment

Tire inflation, strut extension, door and gear actuator rods properly attached, and gene& condition

Nose Flotation Gear s

Condition, security

Pitot Heads

Covers removed, clean, damage, and security

Controliable Searchlight or Landing Light

Lens damage, proper stowage

Emergency Blowdown Bottle

Proper pressure

Electronics Compartment

s closed and security

FAA APPROVED NOVEMBER 21,1978 Revised June 17, 1997

Part 1, Section I1 Normal Procedures FLIGHT MANUAL

FOR -

CHECK Windshield Washer

Fluid level

OAT Source

Clean, free of obstructions, security

Main Rotor

Free to turn; walk rotor through 90" to next 45" position. as necessary

FAA APPROVED NOVEMBER 2 1,1978 Revised June 17, 1997

FLIGHT MANUAL

Part 1, Section 11 Normal Procedures

IN'IXRIOR CHECK The takeoff and anticipation landing gross weight and balance should be determined before takeoff and checked against the LOADING INFORMATION contained in Part 2 and OPERATING LIMITATIONS. NOTE Throughout the remainder of Section 11, checks marked with a plus symbol (+) are required only before the first flight of each day. All other checks should be done before each flight. 1. Cabin exits - Secure and accessible. 2. Cargo - Secure. +3. F i e extinguishers - Charged and secure. +4. First aid kit - Sealed and secure. +5. Emergency locator transmitter Secure, antennas connected, switch to ARM. +6. All other emergency and safety equipment - Condition and secure. 7. Required aircraft documents checked. 8. In-Flight power assurance - Note most current 30-minutepower assurance and daily trend check data. (See IN-FLIGHT POWER ASSURANCE this section).

-

NIGHT FLIGHTS 1. All interior lights - Check. 2. All exterior lights - Check. BEFORE STARTING ENGINES 1. Aft doors - Secure. Check lockpis. 2. Seat and tail rotor pedals - Adjust. 3. Wheel brakes - Set. 4. Safety belt and shoulder hamess - Check and fasten. 5. Continuous ignition switch - OFF.

NOTE Continuous ignition is disabled when engine lever is OFF. 6. Landing gear handle - DOWN. 7. Landing gear emergency T-handle - In.

8. Avionics - OFF. 9. Power on VIE placard - Select. 10. Flight control servo switches - Centered. 11. Generator switches - ON.

-

12. Inverter switch(es) ON.

13. Master start switch - ON.

FAA APPROVED NOVEMBER 21,1978 Revised September 7, 1990

I

Part 1, Section II Normal Procedures

&#[email protected]~~E FUGHT MANUAL

14. Stick trim switches - ON. a. Cyclic stick trim switch - AUTO for AFCS Phase llI.

-

b. Yaw trim switch ON for AFCS Phase llI.

15. Arming Switches - OFF. 16. Channel monitor switch - PBA (if installed). Check both indicator needles centered. 17. Remaining center console switches - As desired. 18. Battery switch - ON. ,

19. Rotor brake switch - ON. Check ROTOR BRAKE caution light on to confirm functioning of 15 psi pressure switch.

If manual rotor brake installed, handle - ON (two strokes required) and check ROTOR BRAKE caution light on and pressure gage for at least 200 psi. 19A. Rotor brake off starts - If desired. With electric trrake installed:

I

a. Select OFF without going to REL, and check caution light off within 3 seconds. b. Select REL, then OFF. With manual brake installed:

I

a. Release brake, handle in detent, and check caution light off within 3 seconds. 20. External power switch (If external power available) - ON

21. Landing gear down and locked lights (3 green) - ON. 22. AFCS switches - OFF. NO GO lights off. 23. Fuel quantity - Check. +24. Fuel indicating system - Note fuel quantity; then press test button. Indicator pointer should decrease slowly to zero. Release test button and pointer should retum to original indication.

25. Cautionladvisory - Press to test. Check these lights: Caution Master warning s T5 repeaters Landing gear AFCS Gomo GO Fire extinguisher test Float test AFCS AIS (AFCS Phase III installed)

FAA APPROVED NOVEMBER 21,1978 Revised July 15, 1999

SA 4047-76-1

~///K@EHKo~&, FLIGHT MANUAL

Pan 1, Section II Normal Procedures

-26. Engine failure warning system - Check. a. # l and #2 ENGINE OUT warning lights - 9 n . b. Engine out ground test switch (side of console) - Hold in TEST. Ahernating tone will be heard. Press ENGINE OUT waming light capsules to reset tone. Warning lights will remain on. c. Engine out ground test switch - OFF. 2'7. Circuit breaker s - Check. 28. Fire extinguisher test. fire detector, prime switch - All switches OFF or XORM. -29. Fire detector system - Check. a. Fire detector test switch - FWD. FIRE warning lights and T-handle lights should light and continuous tone should be heard. b. Press FIRE warning light capsule(s) to reset tone. c. Fire detector test switch - NORM. All FIRE waming lights off. d. Fire detector test switch - AFTIBAG. FIRE warning lights and T-handle lights should light and continuous tone should be heard.

NOTE If external power is in use, the SMOKE DET BAGGAGE warning light will go on.

I

e. Press FIRE warning lizht capsule(s) to reset tone. f. Fire detector test switch - NORM. All FIRE warning lights off. 30. Heater BLEED AIR switch or ECU CONT SEL switch - OFF.

3 1 . Engine levers - OFF 32. Fuel levers - XFEED.

33. Engine T-ha'ndies - Forward. - 3 3 . Engine power assurance target values - power assurance with EAPS ON must be conducted prior to flight with EAPS ON. A check with EAPS ON also assures power with EAPS OFF. Determine t a r s e t T j s e t t i n g s f o r ambient O A T and pressure a l t i tude (Figures 2- 1 through 2- 17). Use Figures 2- 1 through 2-8

FAA APPROVED NOVWBER 21, 1978 Revised May 15. 1992

2-6A

Part 1 Section I1 Normal Procedures

FLIGHT MANUAL

to determine target T5 values f o r helicopters without EAPS o r for those with EAPS installed and with the EAPS switch OFF. Use Figures 2-10 through 2-17 with EAPS installed and with the EAPS switch ON. 35.

Using ambient OAT and pressure altitude, consult the Torque Available, Takeoff Power chart (Figures 4- 6 and 4- 6 ~ to) determine the torque to be used during takeoff.

STARTING ENGINES (BATTERY OR EXTEmAL POWER) NOTE When starting on battery power, the oil press u r e and fuel pressure indicators a r e inoperative until the d. c. generator comes on line; however, the emne oil pressure and fuel press u r e warning lights on the caution-advisory can be monitored for pressure indications. 1.

Rotor a r e a

- Clear

engers and crew should not be allowed to approach o r depart the helicopter directly from the front because of rotor clearance, especially at low r p m and with gusty wind conditions. CAUTICN Rotor brake limits a r e as follows: With rotor stopped, rotor brake operation limited to one o r two engines operating at idle. CAUTION ENGIhTE OPERATION WITH ROTOR STOPPED LIMITED TO: BLADES 45' TO LOXG AXIS UhXESTRICTED

2.

engers and crew

OTHER BLADE POSITIONS NO TAIL WIND COMPONENT AhTD ONLY ONE ENGINE AT IDLE AND ONLY 37" C OAT OR BELOW

- Alerted

FAA APPROVED NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Part 1, Section II SIKDR~IO~ENormal Procedures

SA 4047-76-1

FLIGHT MANUAL

3. KO. 1 engine - Start NOTE Either engine may be started fmt. During external power starts with the battery switch inadvertently OFF, some external power supplies might cause a sharp fluctuation of cockpit indicators. Turn battery on to eliminate. Select freon air conditioner (if equipped) OFF prior to engine starter engagement. Continuous low fuel pressure and the failure to attain light off (with functioning ignition) durin,a a start attempt are indications of possible loss of fuel pressure. l h s condition is more likely to occur following fuel system maintenance or storag of the aircraft. See FUEL PRIMING PROCEDURES in the OPTIONAL PROCEDbRES at the end of Section II for details. CAUTION If rotor should besin to turn during start with the rotor brake on. shut down the engine(s) or release the rotor brake. See Section 111, EMERGENCY PROCEDURES for details.

FAA APPROVED NOVEMBER 2 1, 1978 Revised June 17, 1997

Part 1, Section 11 ANKBRHIO~~E Normal Procedures FLIGHT MANUAL

a. No 1 engine lever - OFF. b. Starter button - PRESS. To minimize possibility of hot starts, or hung starts at lower ambient temperatures, motor engine until T5 is below 150°C and N1 at 12%. c. Engine lever - IDLE. NOTE It is recommended that starting T5 be maintained in the range of 760 to 820°C when possible. At colder ambients, it may not be possible to reach 760°C; therefore, the engine lever should be advanced rapidly to ground idle as soon as 12% NIis reached with Ts below 150°C. At warmer ambients, or when restarting a warm engine, a modulating technique can be used to maintain TSbetween 760 and 820°C. T5 during start should not go over 826°C steady state (transient operation between 826°C and 927°C limited to 10 seconds, operation at 927°C is limited to 1 second). d. Starter button - Release at 58% NI. e. Oil pressure - Above 50 PSI.

FAA APPROVED NOVEMBER 21,1978 Revised March 2, 1989

Part 1, Section II Normal Fkocedures

=&+Rti@~~~flfl& FUGHT MANUAL

f. A normal start should be completed in 25 - 60 seconds. Idle speed is 60% to 61% Ni. NOTE Starter use limits vary depending on whether or not fuel ignition (light-off) occurs in the engine. With fuel ignition the starter limits are: a 60-second start attempt, a 60-second interval, and then a 60-second start attempt followed by a 15-minute interval. Repeat cycle. W~thoutfuel ignition the limits are a 30-second engagement, a 30-second interval, a 30-second engagement, followed by a 15-minute interval. Repeat cycle. g. At idle, check N1 and Ts for normal indications and #1 DC GEN caution light off. NOTE

If the engine has been shut down for more than 15 minutes, stabilize at idle speed for 1 minute before increasing power. CAUTION To avoid auto- (spontaneous) acceleration of the 250€30 or 250-C30S engine, make a ground warmup at idle speed for a period of 10 minutes before flight if all of these conditions are encountered. 1. Low ambient temperatures, - 12OC (10°F) or lower. 2. High relative humidity, 45% or higher. 3. Helicopter has been allowed to cold soak.

If auto-acceleration should occur,shut down engine. Subsequently, restart and resume warmup period. 4. For rotor brake off starts, check transmission oil pressure, servo pressure and tip path response to cyclic inputs. NOTE The rotor brake will hold with both engines at idle, but the following procedures allow the EAPS checks to be most easily done. For helicopters without EAPS, rotor engagement can be made by releasing the rotor brake with one or both engines at ground idle, or without use of the rotor brake as desired. 5. Rotor brake switch (if rotor bxake applied) - Carefully select OFF and check caution light off within 3 seconds. If manual brake installed, move handle to OFF detent and check caution light off within 3 seconds.

I 2-8

FAA APPROVED NOVEMBER 21, 1978 Revised July 15, 1999

Pan 1 Section 11 A~~K@BsK#~~&Normal Procedures FLIGHT MANUAL

NOTE The rotor brake will release as the rotor brake switch is moved to OFF.

6. Rotor brake switch - REL; then OFF. 7. Transmission oil and servo hydraulic system pressures - Check indicators for normal range and MA13 XMSK OIL PRESS warning light and # I SERVO SYSTEM and $2 SERVO SYSTEM caution lights - Off.

7a. No. 1 engine lever - Move forward towards FLY while positioning cyclic to minimize droop stop interference.

8. Flight controls and servo systems - Check at 65 - 70% Nr with droop stops out.

If flight controls do not respond correctly or a restriction in control motion is evident during operation on either isolated hydraulic system, move servo switch to center (both Or\;) and proceed with shutdown.

Do not move flight control servo switch from ON during flight, except in case of emergency where it is necessary to turn off a malfunctioning servo stage. a. Move all flight controls through a displacement of at least 1 inch from trimmed position, checking for binding or roughness. Observe tip-path plane for proper response. b. Pilot's flight control servo switch - NO. 1 OFF. Note that #1 SERVO SYS caution light goes on immediately. There should be no jump in controls. Check flight controls as above.

FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

I

FUGHT MANUAL

Part 1, Section I1 hiorma1 Procedures . .

c. Pilot's flight control servo switch - Move rapidly to NO. 2 OFF. rr"1 SERVO SYS caution light goes off and $2 SERVO SYS caution light sees on immediately. There should be no jump in controls. Check flight controls as above. d. Pilot's flight control servo switch - ON. Both servo pressures normal and both caution lights off. 9. Deleted. 10. No. 1 engine Iever - Advance slowly forward to FLY. 11. AC generator (if installed) - Check ON. AC GEN caution and INV ON

advisory lizhts off. 12. EAPS check (EAPS equipped aircraft). a. Engine anti-ice switches. Check off. b. EAPS switch - Oh'. EAPS caution light will go on, indicating that EAPS doors on No. 1 engine remain open.

13. No. 2 engine - Start. Follow same procedure as for No. 1 engine and then advance slowly forward to FLY. 14. Deleted.

15. EAPS check (EAPS equipped aircraft).

FAA APPROVED NOVEMBER 21, 1978 Revised May IS, 1992

P a r t I Section II Xormai Procedures FLIGHT MANUAL a.

EAPS caution light will go out and EAPS advisory light will go on, indicating that the EAPS is functional. This, in conjunction with steps 12 and 13 above, a s s u r e s that the EAPS automatic door opening circuit is operational.

b.

EAPS switch - CFF. EAPS caution light will remain off and EAPS advisory light will go off. NOTE No. 2 engine EAPS is checked during shutdown.

16. Engine anti-ice

- ON below t 2 " C witln visible moisture.

XOTE With EAPS equipped aircraft, turning on engine antii c e will energize the EAPS system. EAPS operation not permitted in falling and blowing snow. WOTE With the snow protection kit installed, turning on the engine anti-ice will put on the anti-ice caution lights. These lights will go off when the heating elements in the firewall and encjne bellmouth reach about 18°C. 17. EAPS

- As desired. Do not turn the anti-ice/EAPS switches off if accurnulations of ice o r snow is suspected. Opening the EAPS doors under these conditions could r e s u l t in engine damage o r flameout.

18. Zeater o r ECG 19. Avionics

- As desired.

- As desired.

20. Fuel totalizer indicator cator.

- Set total fuel indicated on fuel quantity indiXCTE

The fuel totalizer shall be used f o r information purposes only. The fuel quantity indicating system must b e used f o r all operations.

-

21. F i r e detector test switch AFT/BAG. The SMOKE DET BAGGAGE warning Light on the caution/advisory should Light. Also, f i r e warning Lights and T-handle lights should light and continuous tone should be heard. Return f i r e detector t e s t switch to WORM. All f i r e lights will go off. NOTE

If a f i r e should occur in the baggage compartment, only go on. the SMOKE DET B A G G A G E w a r n ~ l i g h will t (The continuous tone will not be heard.) FAA APPROVED NOVEMBER 21, 1978 Revised July 26, 1982

Part 1, Section II Nonnal Procedures FUGHT MANUAL

22. Fuel levers - DIRECT.

23. External power - OFF and disconnected. 24. Deleted. -25. AC generator - Check at 100% Nr.

a. AC generator overvolt-undervolt test switch - Hold at UNDERVOLT for about 5 seconds. b. AC generator should drop off the line as indicated by the lightins of the AC GEN caution light. Check that JSV ON advisory light o e s on when generator drops off the line.

c. Place AC generator switch OFF. then ON7 to restore generator output. d. AC senerator overvolt-undervolt test switch - Hold at OVERVOLT. Generator dropout should be immediate. Repeat steps b. and c.

,

e. AC generator feeder fault test switch - FEEDER FAULT. Repeat steps b. and c.

+ 26. DC generators - Check. a. BAIT switch - OFF. B A I T OFF caution light - ON. b. NO. 1 generator test switch - Hold at GND FAULT for about 5 seconds.

c. No. 1 generator should drop off the line as indicated by lighting of the # 1 DC GEN caution light. d. Observe that the DC ESS BUS. NO. 1 DC PlU BUS. and the NO. 2 DC PRT BUS remain powered by noting continued fuel quantity and engine torque indications. e. Place NO. 1 senerator switch OFF, wait for 5 seconds, move to RESET to restore senerator output. f. $1 DC GEN caution light - OFF when generator output is restored. Repeat tesr with No. 1 generator test switch at OVERVOLT. Generator dropout should be immediate. g. Repeat steps b. throuzh f. for NO. 2 generator.

h. BATT switch - ON, BATT OFF caution light - OFF

FAA APPROVED NOVEMBER 21, 1978 Revised June 17, 1997

Part 1, Section II N o d Procedures FLIGHT MANUAL

27. AFCS - Check.

NOTE When operating aboard ship it is recommended that the automatic flight control system not be tested and that the system remain disengaged until just prior to takeoff. a. Check all NO GO lights (AFCS control ) - OFF b. Channel monitor switch (AFCS indicator ) - PITCH. c. No. 1 pitch channel engaze-disengage switch - ENG. CH 1 indicator pointer should be approximately centered. d. No. 2 pitch channel engaze-disengage switch - ENG. CH 2 indicator pointer should be approximately centered. Slizht movement of collective pitch should cause both needles to respond. e. Channel monitor switch - ROLL. Engage roll axis following same procedure as for pitch axis (steps c. and d. above). Use lateral cyclic movement to check roll AFCS response. f. Channel monitor switch - YAW. Engaze yaw axis following same procedure as for pitch axis (steps c. and d. above). Use pedal movement to check yaw AFCS response. g. Channel monitor switch - PBA (if installed). Check both indicator needles are full left (LFTEWD). 27A. AFCS Phase I D auto shutdown - Check.

a. Cyclic auto shutdown test switch - LTIFWD. Check for slizht cyclic movement left and forward and that AUTO SHUT - DN caution light goes on. b. Cyclic stick lrim switch - Cycle to ON for 1 second and back to AUTO.

Yaw uim switch - Cycle to OFF for 1 second and back to OK. Check ALTO SHUT - DN caution liht off when the yaw m m switch returned to ON.

c. Cyclic auto shutdown test switch - RT/AFT. Check as in a. and b. above.

27B. AFCS A/S advisory light - Check off.

FAA APPROVED NOVEMBER 21, 1978 Revised June 17, 1997

A FLIGHT l ~ mMANUAL ~ ~ f l f l ENormal ~ Part 1, Section LI Procedures +28. Droop compensator amplifier - Check at 100% Nr.

a. Both engine levers - FLY. b. No. 1 engine - Minimum beep. c. No. 2 engine - Beep forward to split away No. 1 engine N2 about 2% N2Nr (unloaded engine N2 about 100%). Check No. 1 engine below 75% NI. d. Collective pitch lever - Move slightly and rapidly (?1/8") (3mm) up and down. Check that the unloaded engine N2 momentarily increases to the other N m r . Note that the engine driving the rotor will respond with a noticeable power increase ( N I , Ts, fuel pressure, torque and N2 will increase somewhat). e. No. 2 engine - Minimum beep. f. No. 1 engine - Beep forward to split away No. 2 engine N2 as in step c. g. Repeat step d. for No. 2 engine. h. Match torques, set Nt as desired.

NOTE Use tip path plain to control taxi speed. Use wheel brakes only as needed to slow down, turn, stop, or maintain a ground position. 1. Pilot door(s) - Secured. Check lockpins. DOOR OPEN caution li,oht off.

2. Chocks - Removed.

3. Parking brake - OFF. 4. Wheel brakes - Checked.

1. Engine and transmission instruments - Normal range. 2. Caution/advisory - Check.

3. Engine levers - FLY. 4. Heater BLEED AIR switch or ECU CONT SEL switch - OFF.

4A. Heater blower on, or pilot's window open or overhead vents open for ventilation. FAA APPROVED NOVEMBER 2 1,1978 Revised October 19,2001

2- 1OC

%K~RWR~!EQ FLIGHT MANUAL

Part 1, Section I1 Normal Procedures

4B. With EAPS installed - Switch as required.

5. AFCS - As desired (Engage for IMC flight).

6. Stick trims - As desired (Cyclic trim - ON or AUTO for IMC flight and yaw trim as desired).

7. Engine anti-ice switches - As Required. NOTE Engine anti-ice must be on at or below 2°C with visible moisture. 7A. Continuous ignition switch - As desired. NOTE Use of continuous ignition is recommended during takeoff and landing in falling and blowing snow and flight into inadvertent icing, heavy rain, or heavy snow. Discretion should be used during cruise flight due to the reduction in spark igniter plug service life caused by continuous duty. +8. Engine power assurance - Check each engine separately at 100%Nr. NOTE Two engine check procedures are provided. Either procedure should be used at least once daily to ensure the availability of maximum engine power necessary to meet the performance data contained in this Rotorcraft Flight Manual. If acceptance criteria of either procedure is not met, engine maintenance action must be done in accordance with the Maintenance Manuals prior to further flight. The power assurance check defined in steps 8.a., 8.b., 8.c., and accomplished using Installed Power Assurance Check Charts in Figure 2-1 throuzh Figure 2-17, should be used if the pilot wishes to check engine power prior to takeoff. The In-Flight Power CheckITrend Analysis procedure is an alternate engine power check that may be used in lieu of the ground Power Assurance check for those who wish to maintain engine trend analysis to monitor power output. This procedure is defined in steps 1 through 8, under 30-Minute Power Assurance Check, and 1 through 7, under Daily Trend Checks-Normal Cruise, in the cruise procedure, and is accomplished using charts in Figures 2-18 through 2-2 1.

FAA APPROVED NOVEMBER 2 1,1978 Revised June 17, 1997

Pan 1 , Section II Normal Procedures

FLIGHT MANUAL

NOTE Power assurance with EAPS OK must be conducted prior to flight with E M S ON. A check with EAPS ON also assures power with EAPS OFF. This check should be done only when engine and uansrnission oil temperatures are in the normal operating range (green arc).

FAA APPROVED NOVEMBER 31,1978 Revised June 17, 1997

RIGHT MANUAL NOTE

If the aircraft is cold soaked below -16OC ( d ° F ) a warmap period is required to assure proper operation of the torquemeter system. With both engines driving the rotor system. the warmup period is 7 minutes at -1 6°C (3 OF) increasing a t the rate of 1 minute per each 1 OC (1.8"F)below -16°C (+3OF). a.

Using OAT'C and pressure altitude, select lowest possible Tg. target torque Power Assurance Chart and determine -get

b.

With aircraft nose into wind, shut off DC generSLtOr of engine being checked, adjust torque to target valne while - ' - - ,100%

&law at least 10 seconds after initial power application for T j to s t a b i i e and note observed Tg. Hover is permitted if required. Observe height-velocitg limits (Figure 1-4).

. Xr,

c.

Determine TS margin (number of "G M o a target). (1) If the observed % exceeds the target, there is no margin and engine maintenance must be accomplished in accordance with the Maintenance Manual before flight.

(2) If observed T5is less than target Tg by 6 O or mare, suE6cient Tg margin exists and published performance is assured. (Proceed to step d. ) If Tg margin is less than 6", proceed to step c. (3).

(3) If 5 margin is less tban 6O, perform an adpower as-ce checkusingthe next h i g k r t o r q a e chart,

d. DC generatorsrvitch- ON.

9.

10.

e.

Repeat Steps a. through c. for opposite engine.

f.

Match torques at 100% Xr.

Flight instrrrments

- Check and set.

Navigation and communication equipment

-

- Check and set.

11. Windshield heat As required. 12. Pitot heat As reqaired.

-

FAA APPROVED NOVEMBER 21, 1978 Revised .. MAY 14. 1985

Part 1 Section 11 Normal Procedures

NGHT MANUAL

EAPS NOT INSTALLED OR EAPS 1NSTAU.ED AND SWrrtX IN THE OFF POSmON AIRSPEED -.STATIC Nr= 100% ENGINE BLEED A I b O F F GENERATORdFF

TORQUE = 7W

I

POmff TURBINE INAIR TLIIPERAWRE-T 5.C 740 760 780 800 &?O

FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27. 1919

INSTILLEO POWER ASSURANCE CHECK CHART EARS NOT INSTALLED OR EAPS INSTULED AND SWITCH THE O f f POSITION

-

AIRSPEED ~TATIC ENGINE BLEED NR--OFT

I

IN

Nr= 100%

GPIERATObOFF

TORQUE = 75%

I

F i e 2-2

FAA APPROVED NOVEMBER 21, 1918 Revised AUGUST 21. 1919

2-13

Part 1 Section II

No&

Procedures

~ RIGHT f l m MANUAL ~ g ~ d ~ / &

INSTALLED POWER ASSURANCE CHECK CHART EWS NOT INSTALIED OR EAPS IN'XAUED AND SWITCH 1N

-

THEoFFPOSmoN U R ~ E D STATIC ENaNE BLEED A I b O F F

a

Nr= 106%

GENERATOR-OFF

Figure 2-3 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1919

FUGHT MANUAL

INSTALLED POWER ASSURANCE CHECK CHART UPS NOT INSTALLED OR EAl5 INSTILLED AND SWITCH I N THE OFF POSmON

-

AIRSPEED STATIC ENGINE BLEED AIRIOFF

Nr= 10Wh GENERATOR-OFF

Part 1'Section II Normal Procedures

1

Figare 2 4 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

2

Part I Section II Normal Procedures

SRIGHT ~ mMANUAL ?md&

INSTALLED POWER ASSURANCE CHECK CHART

-

AIRSPEED SIATIC ENGlNE BLaD AIR-OR

)IT= 1-

CENERITOR-OFF

Figore 2-5 FAA APPROVED NOVEMEER 21, 1978 Revised AUGUST 27, 1919

I I

INSTALLED POWER ASSURANCE CHECK

CHART

U95 NOT lNSTAUED OR EIPS I W A U E D AND SmCH IN THE MT POSmON AIRSPEED STATIC NI= 1000k ENGINE BLEED A J b O F F GENERAT-

-

-

II

I

Figare 2-6 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1919

2-

UC

Part 1 Section II Normal Procedures

51~amm~dE FUGHT MANUAL

Figure 2-7 FAA APPROVED NOVEMBER 2 1, 1978

Revised AUGUST 27, 1919

Part 1 Section II Normal Procedures

FLIGHT MANUAL

0

9 h

0 0

9 8

-

~30nuL1*3El=-

0

0

a

w

0

9

0

0

9 m

D 0

0

0

b 0

0

9

K

m

9

9 *

'?

P.

C)

0 0

0

-

0

b 0

?

0

-

9

9

9

Q

0

0

L1

CI

N

cri

8

figare 2-8 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1919

FUGHT MANUAL

Figare 2-9.

Deleted.

FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

,-

Part 1 Section I1

K FLIGHT MANUAL

U P S INSTAUED AND SWITCH IN THE ON POSITION

-

AIRSPEED STATIC ENGINE BLEED AIR-OFF

I

TORQUE = tOgb

Nr= 100% GENERATORaFF

I

FAA APPROVED NOVEMBER 21, f 978 &vised AUGUST 21, 1979

RlGHT MANUAL

IUSTALLED POWER ASSURANCE CHECK CHART EAPSIWS~AUEDAND SWITCH INTHE on ON

TORQUE = 75%

Figure 2-11 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

WGHT MANUAL

INSTALLED POWER ASSURANCE CHECK CHART EAPS INSTALLED AND m C H IN THE ON POSITION

-

AIRSPEED STATIC ENGINE BLEED AIR-OFF

I

N r = 160% GENERATMMFF

TORQUE = 80%

I

OAT .C

Figure 2-12 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

1 2- 143

Part 1 Section I3 Normal Proceciures

SNGHT f i rMANUAL n~~~lE

INSTALLED POWER ASSURANCE CHECK CHART E N S INSTALLED AND SWITCH IN THE ON POrmON ~ ~ E E TS ) T ~ C NT= 1 ENGINE BLEED AlRcOFF GENEIUTDRaFF

-

Figme 2-13 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1919

Sfl~@~~rnd&

Part 1 Sectioo ZI Normal Procedures

NGHT MANUAL

I

1 I

IUSTALLED POWER ASSURANCE CHECKCHART lWSTUL€D AND SWITCH IN THE ON

AIRSPEED ENGINE 5-

-

STATIC AIR-OFF

Nr= 100%

GENERA'TOWFF

TORQUE = 9U%

Figure 2-14 FAA APPROVED NOVEMBER 21, 1978 &vised AUGUST 27, 1979

I I 1

I 2- l4L

EIPS IWSTltLED AND SWrCCH IN THE ON POSmON

-

AIRSPEED STATIC ENGINE BLEED URaOFF

I

N r = 1Wh

GENERP,TMMR

TORQUE = 9596

Figare 2-15 FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

I

FUGHi MANUAL

0

f m

.

.

0

s

. )

8 D 0

a d T 0

s 0

8

N 0

0

-. 0

Figure 2-16

FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

0

sm

=

2-14N

I

Part 1 Section Il Normal Procedures

AFUGHT # ~MANUAL Rs~/I&

Figme 2-17

FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

FLIGHT MANUAL TAKEOFF

CAUTION The nosewheel will not self-center a f t e r lift-off if i t h a s swiveled about 180" just before lift-off. Lf retracted in this position, the nosewheel will jam in the up position and n o r m a l extension will not be possible. Lf nosewheel is swiveled m o r e than 90' sLight forward taxi is recommended before lift-off.

F o r night takeoff r i t h controllable searchlight, adjust lighr in hover so that spot appezrs in fronr. just aabove g l a r e shieici. Leave l i g n ~in this position throughout the takeoff. NOTE

When taking off o v e r water. the floats switch should b e placed at ARVED until airspeed approaches 75 KIM.

CATEGORY "A" TAKEOFF 1.

With a i - c r d t a t 100> Sr anci flat pitch, note the altimeter reading.

2.

Hover at 5-foot wheel height.

3.

Set 'Ur at 100';

4.

Increase collective to obtair. t a e o f f torque determined in F i w r e 4-6 (do not go over tzkeoff rating l i c ~ i r s while ) sinlultaneously accelerati?g for\\-zird t o mzlntain a 5- to 10-foot n n e e l height.

5.

At 35 KIAS. rotate tne helicoyter nose up and maintain 35 KIAS.

6.

At the Critical Decision P o k t (CDP), ar! indicated altitude of 40-feet above grocnci altimeter reaciing ( 5 5 feet r a d a r altimeter), adjust the helicoprer zrtitude to accelerate to best r a t e of climb sueed.

itin in matched torques.

CATEGORY "B" TAKEOFF

1.

Hover zt 5-foot heel height.

2.

Set Xr t o

3.

Accelerate forward while maintaining %feet to 10-feet wheel height,

4.

Increase collective t o obtain takeoff torque determined in F i g r e 4-6. Do not exceed takeoff limits.

5.

A s 45 to 50 KIAS is achieved, raise nose to maintain 52 IUAS and

1005with

n a t c h e d torques.

climb until a l l obstructions a r e cleared. During takeoff and climb, observe height-velocity limits ( F i g r e 2-41.

FAA APPROVED NOVEMBER 2 1, 1978 Revised JUhT 28. 1984

2-146

Part 1 Section I1 Normal Procedures

& $FUGHT x f i BMANUAL ~~1dg6

6. After obstructions are cleared, raise landing gear and continue normal flight. CLIMB 1. Landing gear lever - UP before reaching 130 KIAS.

2. Heater or ECL - As desired.

CRUISE If AFCS Phase I11 is engaged, check that AFCS AIS advisory light is on at airspeeds above 50 KIAS. NOTE (Collins DME-40 equipped aircraft) With computed ground speeds below 80 knots the MIN and KTS functions of the Collins, DME with display dashes only. IN-FLIGHT POWER ASSURANCE Introduction An initial 30-minute power assurance check is to be done at 74 KIAS to determine initial power margin. followed by repeated checks at this power level every 50 engine flight hours or as required by engine maintenance or deterioration or at the option of the pilot. Daily trend checks are to be performed at normal cruise power (694°C) at 135 + 10 KIAS at an altitude of nominal 2000 ft or at some other altitude up to that altitude limited by Vne as desired by the operator. If possible, trending should be made at a consistent altitude within = 1000 ft from the selected altitude. Maintain a record of all power margin and trending data. 30-Minute Power Assurance Check Heater or ECU bleed-air - O f F. With EAPS installed - Switch OX. Altimeter - Set to 29.92 inches or 1013.2 Mb to determine Pressure Altitude. DC GEN on engine being checked - OFF. Beep back N2 on engine not being checked. At 74 KIAS, use Nz beepers and collective pitch to set 100°/o KTand 798" T5 on the engine being checked. If 1040/0 torque is achieved before 798" T5, reduce Ts to next lower cardinal 20°C level (780°, 760°, 740°, etc.). When stabilized, record Tj, torque, OAT, pressure altitude and engine anti-ice on or off. NOTE Indicated torque should not be allowed to exceed 104%. Time at maximum T5 temperature should be minimized. FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

Part 1, Section II Normal Procedures

FLIGHT MANUAL 7. 8. 9. 10.

DC senerators switch - ON for both engines. Repeat Steps 1. thru 7. for opposite engine. Altimeter - Reset to current altimeter setting Resume normal flight. Reduce data after flight. See POST SHUTDOWN checks, this section.

Daily Trend Check - Normal Cruise 1. Heater or ECU bleed-air - OFF. 3. With EAPS installed - Switch ON. 3. Altimeter - Set to 29.92 inches or 1013.2 Mb to determine Pressure Altitude. 4. DC senerator switch - OFF on engine being checked. 5. At I35 = I0 KIAS level flight, use N2 beepers and collective pitch to set 100% 'cr and 694" T5 on ensine being checked. If 104% torque is reached before 694" T5, reduce T5 to the next lower cardinal 20°C T5 level (680°, 660°, 640": etc. ). When stabilized, record T5, torque, OAT, pressure altitude and engine anti-ice on or off. 6. DC generator switch - OK for both engines. 7. Repeat Steps 1. thru 6.for opposite engine. 8. Altimeter - Set to current altimeter setting. 9. Resume normal flight. Reduce data after flight - See POST SHUTDOW checks, this section.

1. Crew and engers - ALERTED 2. Kr - AS desired 3. Landing gear lever - Down below 130 KIAS, three geen lizhts - on 4. Heater BLEED AIR switch or ECU CONT SEL switch - OFF 5. Heater blower on, pilot's window open, or door vents open for ventilation. 6. Ensine anti-ice switches - As required.

WARNNG Do not turn the anti-ice switches off with EAPS installed if accumulation of ice or snow is suspected. Opening the EAPS doors under these conditions could result in engine damage or flameout. NOTE Engine anti-ice must be on at or below 2°C with visible moisture. 6A. Continuous ignition switch - As desired. 7. EAPS swkch - ON if required. FAA APPROVED NOVEMBER 21, 1978 Revised May IS, 1992

I

P u t 1 Section II Normal P r o c e d u r e s

FLIGHT MANUAL APPROACH AVD LAXDING Instrun-en: appro2cn speeds of 70 KIAS to 125 KL4S a r c recommended. 1f -4FCS P h s e III i s operating. instrument a p p r o a c h speeds of 60 KL4S tc: 125 KMS a r e recommended. Possible interference between combinations of H F and VOR frequencies can cause l o s s of VOR s i g a ' r u-her. the EP t r a n s n l i t t e r i s keyed. Do not trznsniit or? H F when ~ x a k i n g a VOR approach until it has been determined that it will not i ~ t e r f e r ewith VOR displays. NOTE When making ar. approach o v e r a a t e r . t h e floats switch should be placed t o ARMED when a i r s p e e d is l e s s than 75 KI-S.

FAA

APPROVED NOVEMBER 21. 1978 Revised JUXE 28. 1984

Part 1, Section LI Normd Procedures

FLIGHT MANUAL

Interference from HF transmissions in the 2.0 to 8.0 MHz band will cause loss of ADF navigation sigal. Do not m s m i t on HF in the 2.0 to 8.0 MHz band when performing an ADF approach. During landing and operations near the ground, avoid artitudes greater than 10"nose up, to reduce possibility of tail to round , especidly during operations at aft CG .

FAA APPROVED NOVEMBER 2 1, 1978 Revised June 17, 1997

Part 1, Section II AJKHRSKO~~ENormal Procedures FUGHT MANUAL

NOTE 0

During high speed running landiqs, ailow 2 minutes of intervening forward flizht time above 50 KIAS or 10 minutes of intervening ground time between maximum performance wheel brake applications to permit brake disc cool-down. Collective to yaw control coupling can be felt in the controls at certain extreme control positions not encountered within the normal flight envelope. See Part 2 Section I, Tail Rotor Flight Controls, for derail information.

0

At the pilot's discretion, the copilot will assist in monitoring aircraft attitude.

0

During rotor turning operations aboard ship it is recommended that the automatic flight control system be disengased.

CATEGORY "A'' APPROACH AND LANDIBG 1. Establish approach to arrive at the landing decision point (LDP). (100-feet above touchdown elevation at 50 KIAS and not more than 750 feet-perminute rate of descent.) 2. Continue the descent to about 50 feet above touchdown, then reduce the rate of descent with a cyclic flare to about 20" nose up.

3. Level the nose to 5" to 10" at about 30 feet above touchdown. Apply collective to establish hover or cushion the touchdown.

4. After landin:,

reduce the collective pitch to minimum.

5. Apply brakes as required. CATEGORY "B" APPROACH AVD LAhSDmG I . Estabiish approach to arrive at a point 100-feet above the touchdown eievation at 50 KLAS at a rate of no more than 500 fpm.

2. Decelerate to 50 feet at 40 KIAS. 3. Continue approach and deceleration to runing touchdown or hover. Observe limitations shown on height-velocity dia=pm (Figre 1-4).

FAA APPROVED NOVEMBER 21, 1978 Revised June 17, 1997

I

GO- AROUND

1. b c r e a s e collective pitch up to takeoff power. 2 . Accelerate to best rate of climb speed.

3. Raise l a n w gear when positive rate of climb is established.

PRE- STETDOWN 1. AFCS switches 2.

- OFF

Continuous ignition switch

3. Parkingbrakes

-

- OFF.

Set

4 . Flight controls - Centered, collective down 3.

Avionics - OFF

SHUTDOWN NOTE If an internal engine fire occurs a s indicated by a rapid increase in Tg, immediately e m u e starter with engine lever and fuel lever OFF to minimize high Tg, and put out fire.

MOTE Droop stops shoulci be in a t about 45% Nr. Tip path plane clearance is reduced if droop srops a r e nor in. NOTE s o r m a l Nr for rotor brake application with one engine shut down and the other e q g e a t IDLE or with both engines shut down is 50% to 60% Nr. Maximum for shutdown is 65%. For emergencies with both e q g e s shut off, roror brake may be applied up to 10100Nr. CAUTION

Maximum manual rotor brake application is not recommended when commenced below 30% Nr. Brake applications begm a t l e s s than 30% Nr should be light enough to produce a smooth, gentle stop. CAUTION ENGINE OPERATION WITX ROTOR STOPPED L I i i E D TO:

FAA APPROVED NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

FLIGHT MANUAL

BLADES 45" TO LONG AXIS OTHER BLADE POSITION .

UNRESTRICTED

NO TAIL WIND COMPONENT AND ONLY ONE ENGINE AT IDLE AND ONLY 37°C OAT OR BELOW

AIRCRAFT WITHOUT EAPS 1. Engine levers

- IDLE (for 2 minutes)

2. Engine leveds)

- OFF

3. Rotor brake switch caution light on

- As desired.

If ON, check ROTOR BRAI(E

-

3 k Manual rotor brake (if installed) hnncne As desired If OX, (two s-es required), check ROTOR BRAKE &on light on 4. Remaining engine lever O F F

-

I

1. Nr- Maintain 100% 2. Engine anti-ice switches

- Check O F F

- IDLE (for 2 minUtes) 1 engine lever - O F F

4, No. 1 engine lever

5. No.

6. When No. 1 engine N1 drops below 58%, EAPS cantian light will go an (after a 5-second delay), mdiNo. 2 engine EAPS doors have opened This will ihat the automatic door opening circuit is operational.

7. No. 2 engine lever

- IDLE (for 2 minutes)

-

If ON. check ROTOR B W 8. Rotor m e switch As .dcaution light on 8 k Manadl rotor brake (if installed) handle As desired. If OX, (two strokes required), check ROTOR BRAKE cavtibn light an

-

9. No. 2 engine l e v e r

- OFF

POST smJTmwN

-

1. lbel levers OFF 2. Center cansole d t e h e s

- As desired

FAA APPROVED NOVEMBER 21; 1978 Revised FEBRUARY 22, 1985

I

Part 1 Section II Normal Procedures

=~~~@RBKo#& FUGHT MANUAL

4.

Battery switch

- OFF after rotor stops NOTE

Turning battery switch OFF before rotor stops can cause rotor brake to revert to high pressure, causing a xapid rotor stop. 30-Minute Power Assuance Calculatior.

1.

Using the data collected during the flight. enter the 74 MAS Installed Power Chart (Figure 2-18 f o r EAPS not instdlled. and Figure 2-20 for EAPS installed and switch OX) at OAT, proceed vertically to .applicable Tg h e , and follow horizontally to intersect the pressure altitude h e ; then drop vertically to read target percentage torque.

Apply anti-ice ON correction factor. if necessary. (Figure 2-22). Repeat for number two engine. a.

b.

c.

If the observed torque f a l l s below the target torque (negative =he). %re is no margin and engine maintenance must be done in accordance with the Maintenance M a d before next flight. E observed torque m2rgi.n is zero, or more, sufficient margin exists to establish availability of power required, and published performance is assured. Record data to include date and engine hours. NOTE Repeat SO-Mirmte Power check: eve- 50 engine flight h w , following engine mainte~=ce.

Daily Trend Check Calculation NOTE For each engine, - ' - a daily record d tarqae margin as determined from Paragraph 3 below and establish a '?orolling" average (trend 3ine) for 20 5 consecutive data points of operation.

-.

I.

Obtain the most recent average e e af t o q u e margin based on the most recent consecutive 20 i 5 data points af operation.

2.

Using the data collected during the flight, enter the 135 gLAS Installed Power Cbart (Figure 2-19 for EAPS not instdlled. and Prgare 2-21 for EAPS installed and m t c h ON) at OAT, proceed vertically to the a p plicable 5 h e and follow holizontduy to the intersection of the pressure altitude line: then drap vertically to read target percentage toque. Apply anti-ice ON corrective factor, if necessary (F2-22) Repeat for number two engine.

2-16B

FAA APPROVED NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Part 1section II Narmal Procedures NGHT MANUAL

3.

Determine torque margin (obsemed torque minnn w e t torque).

a.

.

b.

SxEEicient margia exists to establish of power m e e d and published performance is assured if: (1) Observed torque margin is not more than 4% below aveage torque margin from Step 1. (2) Average torque margin i s zero o r higher. (3) Average torque margin bas not deteriorated by more than tbe margin determined from the last 30-Minute Power Check. If during any 50-hour baseline i n t e n d , the observed torque margin is more thaa4%below the average margin, o r the avexage torque margin falls below zero, or the average torque margin deteriorates by more than the margin determined from the last 30-Minute Power Check., Derform engine maintenance in accordance with the Mah&name Manual b r e the next £light.

-

4.

Record data to include date and engine hours.

(Auti-Ice O W ) established Step 1 of 30-Minrrte Power w Calcn3ationaadStep2of~T~checBCalc_rzlation, FigrPe 2-22.

e

1

Anti-Ice On Correction Factor

EXTERNAL CARGO BOOK OPERA!lXON

NOTE The preceding normal procectares shodld be followed unless superseded by tke following

The following checks should be perfarmed daily, prior to cargo hook apedons: 1. Ground personnel exerts pressme on cargo hook from beneath the

helicopter. FAA APPROVED NOVElVIBER 21, 1978 Revised JANUARY 17, 1983

2-16E

Part I Section I? Normal Procedures

SFUGM N ~MANUAL ? H K # ~ E 4047- 7& 1 SA

-

2.

Mechanical release handle WLL. G r d personnel should confirm hook open. Mechanical release handle rmSH down. Ground personnel should confirm hook locked

3,

Cargo hook arming switch light on.

4.

Pilot's cyclic release button DEPRESS. HOOK OPm advisory light will cycle ON/OFF. Graaad persounel should confirm hook open and allow hook to relock when cyclic button is released.

5.

Copilot's cyclic release button

-

- ARMED.

Check HOOK ARMED green

-

- DEPRESS.

Repeat checks.

When an extiernal load is being carried, the following performance factors should be considered.

I. The helicopter was tested for this operation under day, VFR conditions only. This does not necessarily preclude the o-perator from demonsm%t-gthe ability to operate under night, VFR conditions. 2. Performance published in the £light manual may o r may not be realized depending on the size and s h a p of the load to be carried on the sling. Compensation should be &owed for the d m g effects of ext e n d loads having iarge surface areas. 3,

Figure 4-21 shows maximum gross weigbt for hovering out of ground effect. Vertical drag of Ute external load is not ed for.

2.

Ground personnel-Secure load to cargo hock. Under some atmospheric conchtions, a - a c electricaZ charge w i l l exist on the r o t o r c a f t in night. This charge should be dissipated by a suitable means prior to ground personnel touching the cargo hoak of a hovering rotorcraft.

3.

S h l y apply collective and ascend vezticdlly to apply tension slowly to the cargo cable and to insure that cyclic control is adequate and that power limits are not exceeded.

NOTE

The CG shift with extend load is approximately equal to 2.5 inches per 1000 pounds of load weight i. e. , a 3300pound load w i l l r e d t in a foraard CG shift of approximately 8 inches.

CRUISE Accelerate slowly into forward flight to that no condition is encountered in which the rotorcraft can be uncontrollable. FAA APPROVED NOVEMBER 21, 1978 2-16F Revised JANUARY 17, 1983

FUGHT MANUAL APPROACH AND LANDING

The descent and approach to the externdl load drop area should be made at a combination of horizontal and vertical speed s o that the r o t o r c a f t w i l l come to a stabilized hover w e t the drop point. Descend slowly with minimum fo-d speed and no side drift until the cargo rests on the gramrd and the cargo attachment cable becomes sfack. CAUIIZON Excessive approach speeds o r rates of descent may require the use of excessive torque at termination of the approach. R. H. SrSDING CABIN bOOR (MODTFED BY FLIGHT OPERATION.

P/N 76070-20015-011)IN

WAFtNlNG The F A S m SEAT BELT sign must be an and all cabin occupants shall have safety belts fastened o r

wear an apprwed safety harness, properly adjusted, witb the sliaine: door open in w h t .

Avoid prolonged dawnwind hwer witb the ~ l i d i u g door open to prevent accuxndatian of &anst fumes in the cabin. I.

Cabin occupants

- Seat belts and/or safety harness fastened

Safety harness ehbe attached to an m d cargo tiedown ring. Tiedown ring shotiid not be loaded lakerally when a ring is in the vertical position. Ttte ring should be free to swivel s o the terrsianisappliedtothetopaftherirrgrattterthan the side.

3.

-

R.H, Sliding Eabin door Open fully, The & CAUTION D O O R O P E N c a ~ andRHDRlightsgoan, ,

FAA APPROVED NWEM3ER 21, 1978 Revised JANUARY 17, 1983

I

I

4,

-

Press RBDR Egbt MMTEFt C A U n O N and DOOR OPEN capsule goes off. The REDR light stays on.

NOTE can be a c d x a t e d rapidly, especially in the low-speed range. With the R.H. slic?iagdoor open, cfranges in airspeed cues (noise, M e t i n g , ets) are minimal. Exercise care to prevent exceeding the 74 MAS sfidiag door airspeed limit. The *aft

1

5.

-

Aft door pin eqagement Check upper and lower pins visuaUy. Confirm door is locked open by attempting to dide it Solward.

U T I U l T HOIST OPERATION

Check trtifity hoist hook for operation, swivel for freedom of movement and fairing far security.

Determine that lateral and cg limits w i l l not be exceeded drving planned hoist Qperatiarrs. See Part 2 S e c t . XII. 1. &hmal ~ a b k cU&rs Attached and secure. 2. Wet9 hamess Secare and adjnsted. 3. H ~ i s t ~ ~ - S e c r a e a n d ~ ~ a n n e c t e d

-

-

FAA APPROVED NdVEMBER 2 1, 1978 Revised 3ANUARY 17, 1383

P a r t 1 Section I1 Normal Procedures

FUGHT MANUAL

PREFLIGHT

R

This check requires both dc generators operat-. Hoist high speed w i l l not function with only one dc generator o r external power. Only the hoist shear w i l l function with batiery power. Hoist power switch ON. Green indicator light on. -

-

-

-

L i t guard and place switch to TEST. Cable shear test =itch Observe that hoist p m e r indicator goes out. Pilot's cable shear switch - Lift g u a r d and place to SHEAR (UP). HOIST TEST LT near cable shear test switch and HI TJ3MP WARN light on pendant should light. Pilot's cable shear switch

- Close w r d .

Both lights shouid go off.

Repeat steps 3 and 4 usmg cable shear switch in pendant control storage compartment.

-

Cable shear test switch Close guard. Switch will move ro -4RM Check hoist power indicator green light ON. Hoist master switch pendant conrrol.

-

- CREX.

Check operation of hoist ~ i t h

Hoist master switch - PILOT. Check operation of hoist conuol =itch on control console. HOIST power switch U E U T Y HOIST PImP

- OFF.

- Secured and adjusted.

1.

Hoist operator's safety hamess

2.

NOTE Hoist operator snould put on the safety harness and zttzch it to the floor before opening the sliding door. Harness should be attached to an unused tiedown ring. R. H. Slidiig cabin door - As desired.

3.

EC'L' o r heater blower off. XOTE and With all possible helicopter avionic, light-, ECS/ECU systems ON, a single DC generator failure could cause the re- during hoist operations mammg DC g e ~ e x a b to r exceed i t s rated load. *

- ON.

4.

Hoist power switch

Green indicator light on.

5.

Hoist aster switch - PILOT o r CREW as desired.

FAA APPROVED NOVE-WER 21, 1978 Revised JAhiARY 17, 1983

Part 1, Section I1 N o d Procedures

FLIGHT MANUAL 5A. It is the hoist operator's responsibility to assure that hoist cable does not

any portion of the aircraft. In situations where with airframe or other obstacle is observed, hoisting should be intempted while the area on the cable is inspected. If any broken wire, unraveling, or kinks are observed, hoisting operations should be discontinued and the cable replaced. WARNING

Reeling a kinkedldamaged cable into the hoist may cause a hoist jam condition when reel-out is anernpted, renderins hoist inoperative.

5B. Hoist crewmadoperator is responsible to maintain stability of the hoisted load by use of hoist controls, ICS calls to pilot, and physical control of cable (hand or foot). For minor oscillation (linear or circular swing), stop reel-in, apply hand motion to cable in direction opposite to oscillation. For significant oscillation, stop reel-in, srart reel-out or call for pilot to lower aircraft.

If not quickly stopped, the instability may become unmanageable. Reeling in an unstable load will only aggravate the motion. 5C. All crew should observe for shock loads, jerks, or snaps that impart high

loads on cable. If observed, hoisting should be interrupted and cable inspection undertaken to inte-pity (no broken wires, unraveling, or kinks) before resuming operations. Replace cable before next mission.

-

2- 16K'

Static elecmcity should be dissipated by suitable means before sound personnel touch the hook or cable.

FAA APPROVED NOVEMBER 2 1, 1978 Revised June 17, 1997

Part 1, Section II

~//K@RSKO~~&Normal Procedures FUGHT MANUAL

6. Hoist cable - Lower.

The last 20 feet of cable is painted red and is unusable. 7. Ground personnel - Secure load to utility hoist hook.

8. Hoist cable - Apply tension by either raising cable or increasing hover height. Adjust cable len-gth as desired. NOTE Swinging loads may cause erratic radar altimeter and ADF indications.

The hoist operator should wear gloves to prevent injury from broken wires. Cable with the rails or slidins door in forward flight may be avoided by carring load with at least 100 feet of cable extended when practical. 9. Hoist master power switch - OFF. 10. R.H. Slidins cabin door - As desired.

NOTE With a constant load on the hoist, symmemcal reduction in gross weight, for example, fuel bum off, will result in the lateral cg moving outboard. See Part 2, Section IU. HOIST D L ?CYCLE Hoist motor is protected with a thermal shutdown circuit. To preclude inadvertent shutdown, recommend this duty cycle: 1. Between each cycle (reel-infreel-out),pause 30 seconds. 2. For continuous cycles, rest 30 minutes between each series of 5 cycles for a 600 lb. load, or 8 cycles for a 300 lb. load.

FAA APPROVED NOVEMBER 21,1978 Revised June 17, 1997

Part 1, Section I1 Normal Procedures

FLIGHT MANUAL OPTIONAL PROCEDURES

ENGINE FIRE EXTINGUISHER SYSTEM TEST The fue extinguisher test provides a method of performing a full functional check of the fire extinguisher system from the cockpit. This test may be performed whenever desired by the piiot or as required by maintenance. Make this check before starting engines.

I. Fire exringuisher test switch - OPEN. Check red WARN light on.

2. Pull both T-handles back 1 inch. 3. F i e extinguisher - MAIN, then RESERVE. Check that green TEST light goes on in each position. Release switch. 4. Fire extinguisher test switch - SHORT. Check red WARN light and green test light ON.

-

5. Fire extinguisher switch MAIN,then RESERVE.Check that green TEST light stays on in each position. 6. Both T-handles - Full forward. 7. Fire extinguisher test switch - NORM.

8. Check red WARN light and green TEST light off. FLOTATION SYSTEM TEST Make this check for electrical continuity daily before all anticipated over-water flights. DC external power must be available or the helicopters dc generators must be operating to make this test. 1. Flotation test switch - TEST. Red

2. Floats arming switch Check on.

WARN indicator light - Check on.

- ARMED. FLOTATION ARMED advisory light -

3. Floats inflation switch (piiot's cyclic) - Press. Check TEST 1 and TEST 2 n-, indicator lights - On. Release switch, lights out.

-

4. Floats inflation switch (copilot's cyclic) Ress. Check TEST 1 and TEST 2

,oreen indicator lights - On. Release switch, lights out.

FAA APPROVED NOVEMBER 2 1. 1978 Revised March 16, 1994

~//R@RHKW?& FUGHT MANUAL

Part 1, Section II Normal Rocedures

5. Floats arming switch - OFF. FLOTATION ARMED advisory light - Check off.

6. Flotation test switch - NORM. Check red WARN indicator light - Off.

FUEL PRiMING PROCEDURES - DURING NORMAL START Loss of fuel prime can be indicated by continuous low fuel pressure and the failure to attain light off (with functioning ignition) during the start attempt. This condition is more likely to occur following engine or fuel system maintenance or prolonged storage of the aircraft. The following prime procedure should be used to attain light off. If loss of prime is noted after relatively short intervals, suspect a malfunctioning fuel line check valve.

FAA APPROVED NOVEMBER 21,1978 Revised March 16, 1994

I 1

Part 1, Section I1 ~IR@RHM~E Normal Rocedures FLIGHT MANUAL

A. Starter motoring: In a sea level environment, two smt attempts with the fuel lever in DIRECT (30 seconds start, 30 seconds off, 30 seconds start) will usually result in a successful light off by the end of the second start attempt.

B. Engine cross prime: With one engine running, the other engine can be primed follows: Continuous ignition switch - OFF Fuel lever of engine to be primed - PRIME Engine prime switch of engine to be primed - PRIME Fuel pressure gage (if installed) - Note fuel pressure rise Engine - start Engine prime switch - OFF Fuel lever - crossfeed. NOTE Extended running of an engine in D I R E n with the opposite engine's prime switch at ON and fuel lever at PRIME or XFEED can result in vent overflow if tanks are full. C. Manual primin?: The engine can be primed after maintenance using the fuel prime port in the engine fuel system. Consult the Maintenance Manual for details. NAV TEST SWITCH FUNCTION

(When used in conjunction with King Radio Navigation System) When NAV TEST switch is at PILOT or COPILOT, all of the correspondne pilot's or copilot's Marker Lamps should go on. If the appropriate Marker audio selector switch on the ICS conuol is selected, a tone should be heard. (When used in conjunction with the Collins Radio Nav. Sys.) 1.

Turn on VOR receiver(s) and select a VOR frequency on the appropriate Nav. Control . Set pilot's HSI display switch to PLT NAV; copilot's HSI display switch to LT NAV; HSI OBS switch to NORM; PLTILT bearing switches to VOR.

2. Set and hold NAV TEST switch at PL'T (to test pilot's system), or LT (to test copilot's system).

3. Select the corresponding HSI Course Select Knob to center CDI.

FAA APPROVED NOVEMBER 2 1, 1978 Revised May 15, 1992

Part 1 Section I1 Normal Procedures

AFUGHT # KMANUAL @ R- H ~ ~ ~

The following should occur on the appropriate indicator: a. The NAV and Both G/s Warning Flags should remain out of view. b. The deviation bar should center at a course of 000' +5' and G/s deviation bar should deflect approximately 1-1/2 dots down. c.

The ~ o / ~ r oindicator m will indicate TO.

d. The bearing pointer will indicate between 000' 4'. e. All of the pilot's (or copilot's if LT position is selected) MKR BCN lights should illuminate. Release the NAV TEST switch. Select a Locdlizer frequency on the appropriate nav control , and rotate Course Set Knob to select the aircraft indicated heading. Set and hold the NAV TEST switch in PLT position (to test pilot's system) or LT position (to test copilot's system). The following should occur on the appropriate indicator: a. The LOC and Both G/S Warning flags should remain out of view. b. The Localizer deviation bar should deflect approximately 1- 1/2

dots to the right. c. The G/S deviation bar should deflect approximately 1-1/2 dots down. d. All of the pilot's (or copilot's if LT position is selected) Marker Lamps should illuminate. Select MKR audio on the appropriate ICS control. A tone should be heard. Release the NAV TEST Switch. FUNCTIONAL CHECK 76O763OOO6012

- SNOW PROTECTION KIT P/K 76080-30008-011

or

This full functional check of the snow protection kit must be pedormed daily when flight in fallix o r blowing snow is anticipated. It is to be done with both engines running, both dc generators and the ac generator operating.

-

1. No. 1and No. 2 engine anti-ice switches OX. Check that +l and 62 ENG ANTI-ICE caution lights and PI and P2 ENG &'TI-ICE OK advisory lights go o n FAA APPROVED NOVEMBER 2 1, 1978 Revised July 26, 1982

Part 1, Section II Normal Procedures -.. -

FUGHT MANUAL NOTE If #I and #2 ENGINE ANTI-ICE caution lights do not go on, there is either a snow protection system malfunction or the ambient air temperature is too warm (about 18°C OAT) to permit a valid check of the snow protection system. 2. Check that the #I and #2 ENG ANTI-ICE caution lights go off within 2 minutes and remain off, and that the #I and #2 ENG ANTI-ICE ON advisory lights remain on. 3. Converter circuit breaker (on auxiliary circuit breaker ) for 7608030008-012, or engine #I and #2 inlet anti-ice 10 amp ac circuit breakers (below lower center circuit breaker ) for 76076-30006412 - PULL.

- OFF. Check that #I and #2 ENG ANTI-ICE caution lights go on within 4 minutes, and that # 1 and #2 ENG ANTI-ICE ON advisory lights remain on.

4. No. 1 generator switch

NOTE In case of failure of the snow protection kit in a d snow conditions, the caution lights will go on in a considerably shorter period of time.

-

5. No. 1 generator switch ON. Check that # 1 and #2 ENG ANTI-ICE caution lights go off within 2 minutes.

6. No. 2 generaor switch - OFF. Check that #1 and #2 ENG ANTI-ICE caution lights go on within 4 minutes, and that #I and #2 ENG ANTI-ICE ON advisory lights remain on.

7. Reset converter or engine #I and #2 inlet anti-ice circuit breakers. Check that #1 and #2 ENG ANTI-ICE caution lights go off within 2 minutes. 8. No. 2 generator switch - ON.

FUNCTIONAL CHECK - PITCH BIAS ACTWATOR During nominal 100 KlAS cruise, note that both PBA indicator needles are approximately centered. The needles should track together (upper needle motion somewhat delayed), and should move to the right with increasing airspeed and to the left with increasing collective pitch setting. NOTE

Do not pull the BIAS ACTR circuit breaker for failure of the PBA actuator to move. FAA APPROVED NOVEMBER 21,1978 Revised March 16, 1994

1

SECTION III EMERGENCY PROCEDURES

The procedures outlined in this section deal with the common types of emergencies; however, the actions taken in each actual emergency must relate to the complete situation. Extraordinary circumstances such a s compound emergencies may require departures from the normal corrective procedures used for any specific emergency. Throughout this section, the 'land immediately7', 'land as soon as possibler' and 'land as soon as practicable" a r e used to reflect the degree of urgency wiih which a landing must be made. Land immediately

- Self-explanatory.

-

Land a s soon a s ~ o s s i b l e Land at the nearest site at which a safe landing can be made. Land a s soon a s practicable - Extended flight is not recommended. The landing site and duration of the night a r e at the discretion of the pilot. Many of the malfunctions described in this section w i l l be indicated by the lighting of warning o r caution lights, the master caution light, and in some cases, a tone in the headsets. Whenever a caution light goes on, the master cauuon liet capsule should be pressed in to put the master caution. light off, and reset it for another condition. An audio tone can be eliminated and reset for another condition by pressing the appropriate warning light capsule. NOTE Several of the following procedures require pulling and resetting circuit breakers. The circuii: breakers referenced in these procedures a r e marked in white for ease of identificauon. When locating circuit breakers at night, it is recommended that the glare shield lights be set to BRIGHT. EXGmE SYSTEM MAL FVTCTIONS En,oine system malfunctions generally a r e of three types: (1) partial o r totdl failure of the engine to provide power as demanded, (2) failure of an engine control system, (3) failure of engine auxiliary systems, (lubricating system, torque system, N1 indication, etc.)

In case of any malfunction, the f i s t duty of the crew is to make sure of continued safe flight until the malfunction can be diagnosed. Particular care must be taken to confirm which en,@ne has the malfunction. SmGLE-EXGINE POWER FAILURE Symptoms: Initial: Large torque split

FAA APPROVED NOVEMBER 21, 1978 Revised MARCH 16, 1979

Part I. Section I11 Emergency Procedures

Confirrnin,o:

&#fz%~g~g6 FUGHT MANUAL

Nr remains the same or droops NI decreases below 5 9 8 T5 decreasing # 1 or #2 ENG OUT warning light and alternating tone

SINGLE-ENGINE POWER FAILURE WITH EAPS OPERATING OR AhiI-ICE OPERATING With anti-ice off and EAPS operating, EAPS for both engines will be automatically shut off (doors open) if one engine fails. This will maximize power available from the operating engine. The EAPS caution light will o on and the EAPS advisory lisht will go out in this case. With anti-ice on, both anti-ice and EAPS (if installed) will continue to operate for the remaining en,oine. The ANTI-ICE caution light for the inoperative ensine will go on and the ANTI-ICE ON advisory light for the inoperative engine will go off.

I

Do not turn the anti-ice switches off with EAPS installed if accumulations of ice or snow is suspected. Opening the EAPS doors under these conditions could result in engine damage or flameout. SINGLE-ENGINE FAILURE - HOVER (5 TO IS FEET) 1. Maintain collective pitch setting or lower collective slightly if required to

establish descent. 2. Increase collective to cushion landing as touchdown becomes imminent.

3. After touchdown, neutralize cyclic and simultaneously reduce collective.to minimum.

SINGLE-ENGINE FAILURE ON TAKEOFF - CATEGORY A Engine Failure during Category A takeoff can be detected by the occurrance of an unusual torque split and possible rotor droop. Exact procedures to be followed depend upon the point in the takeoff sequence where the failure occurs. Failure before or at the Critical Decision Point (CDP) dictates that the takeoff be rejected. If the failure occurs after the CDP, the helicopter can be flown away from the heliport. During the cat ego^ A climbouts, 96% Nr is used to maximize ciimb rate for speeds from Vtoss (52 EUAS) to best rate of climb speed. For single-ensine flight above best rate of climb speed, use 100% Nr or higher. CAUTION During high speed running landings, allow 2 minutes of intervening forward flight time above 50 KIAS, or 10 minutes of intervening gound time between maximum performance wheel brake applications to permit brake disc cooldown. FAA APPROVED NOVEMBER 21, 1978 Revised May 15, 1992

Hover

1.

-

Refer to pamgraph titled, SINGLE-ENGINE FAILURE HOVER (5 TO 15 FEET).

Prior to Initial Rotation Speed d 35 KIAS 1.

AdfPst nose attitnde to 5 * to lo0 nose-up .

2.

shion to touchdown with collective.

3.

After tachdown, lower collective completely, centering cyclic as collective is lowered.

4.

Apply brakes as reqnired to stop within heliport confines.

After Initial Ro!ation During Climb up to and Including the CDP (35 KIAS and 40 Feet above Ground Altimeter R(55 Feet Radar Altimeter)). 1

Lower collective positively to contain Nr and prevent extending landback unnecessarily.

3.

As helicopter apprcaches heliport, redoee ate of descent with a cyclic fbre to about 20" nose up.

4.

Level nose aUilude to So to lo0 nose up at about 30-feet abom= touch-

5. 6.

down. Apply collective to cushion touchdown. After touchdown, nerrtrdtize cyclic and sinmlhneously reduce colleetive to minimum. Apply b a k e as required to stop raithin codines of -rt.

After CDP

I.

L

3.

M j c s t collective to maintain engine Tg and t o m e

4. 5.

6.

m cdlective to should be mainbbed.

Nr arithin limits. 96% b 100% Nr

at 2.5 minrtte power limits (826O Tg and/or 111% t o m e ) . Mainbin limits at all times. Use N2 beeper to set 96% Nr. Keep airspeed a t o r below best ate af dlimb speedwhile operatingat 96% to99%Nr. Landing gear up when positive rate of climb is established. When cleazaace is assured, accelerate to best rate-ofclimb speed. Cantinne to climb at g6% Nr until crrdse altitade be maintained, then adjast Nr to 100% to 107%(as desired) with N2 beeper and cmtirme flight.

FAA APPROVED NOVEMBER 21, 1978 Revised MAY 14. 1985

-

FAIl,UREON TAKEOFF CATEGORY rrB"

mGLE-EN-

If gross weight and flim path permit, takeoff and climbout m a y be contimed. For a rejected takeoff, do this:

-

I. Collective pitch Reduce as necessary to maintab rotor rpm if altifade permits. 2. Make a partial flare at about 50 feet to minimize ground speed.

-

3. Cyclic and collective Apply to cushion gmtmd cantact

I

4. After touchdown, neutmlize cycUc and tive to minimum.

',

"

arrsly reduce collec-

5. Apply wheel brakes to minimize ground roll.

I. Adjust dl-e as necessary t~ - ' - desired rotor rpm and toque. If reqrrized minimize altitde loss by applying up to maximmn OEI p e r consistent with engine and rotor limits.

3. Cmsider air restart. Fallow procedures in p a a g a p h titled Engine Restart in Flight. Restart should be attempted oaly if the cause of the initial failure has been determined and corrected 4. If restart fails or no attempt to restart is made becaase of coaditians causing engine failure, follow procednre in paragraph titled ~

S

h

I

t

t

d

~

i

n

~

t

FAA APPfPDVED NOVEMBER 21, 1978 Revised MARCH 23, 1981

Engine Restart in Flight An engine restart may be attempted after a flameout subject to the pilot's evaluation of the cause of flameout. The following procedure assumes &at the affected engine was operating with its fuel lever in DIRECT. If a flameout occurs with the fael lever in XFEED, attempt restart with the fuel lever in DIRECT (no prime will be available).

Increased generator load during an air restart could cause a significant Tg increase on the operating engine. The power setting on the operating engine should be adjusted to allow this temperatnre increase without exceeding 826"C Tg. T-handle (affected engine) F o m d .

-

2. Engine lever (affected engine) 3. FLel lever (affected engine) 4. Continuous ignition switch

- OFF.

- XFEED.

- OFF. -

Engine prime switch (afEected engine) OX. 6. Master start switch Oh'.

5.

-

-

7. Operatisg prrpine R-ce output power with collective pitch to maximum contirmous power, 768°C Tg, or below. 8. Engine

- Start (as in normal procedure).

9. Engine running at idle

-

- Engine lever full forward.

10, % and brqoe Adjust as desired. 11, Engine prime switch OFF.

-

Nbm Iftherestaztattemptfailsand/ornofurtherat&mpts are to be made, shut the engine down.

h e 1 Crossfeed After Engine Fgilnte CAUTION Fhel crossfeed after an engine failure should an3y be COPSideredifnecessarytopetmitfiighttoa~le landing area There is same possibriitp that a fnel system malftmctian was tke case of the fust engine failwe, and that crosslmight expose the remainjng engine in a similar problem. The following procedm-e will minimine %e chance of interrapfuel flow to the remaining engine. If rmsteadg operatimoftheoperatingengineiSnotedatanytime&sring the procedure, rapidly discontinue attempts to uossfeed

FAA APPROVED NOVEMBER 21, 1978 Fterised July 26, 1982

1.

/ I

-

Failed engine Shut down for at least 5 minutes to minimize possible ignition source f o r fuel Leaks.

-

engine b e 1 lwer PRIME. Failed engine prime switch PRIME for at least 1 minute, (monitor operating engine fuel pressure, fuel now, N 1 and R. If signs of unsteady operation are noted, quickly move the fuel lever of the failed engine to OFF and discantinne attempts to crossfeed. ) 4. Failed engine fuel lever OFF.

2. 3.

-

-

-

5. Failed engine prime switch OFF. Operating engine fuel lever

-

XF'EED. Monitor operating engine fuel pressuie, fuel flm, Nl, and Tg. If sigas of unsteady operation are noted, quickly switch fael lever back to DIRECT and discontinue attempts to crossfeed. Engine Shutdown in Flight 1. Eb@ne lever OFF. 2. h e 1 lever OFF. SINGLE-ENGINEFAILURE DURING APPROACE.- CATEGORY "AT' 6.

-

-

The Category ""A" Landing Decision Point (LDP)is 100 feet above the touchdown elevation at 50 KLAS and not more than 750 fpm rate of descent. With an engine failure before o r at the I 9 P , the pitot may either conliame go-around. With the approach to a touchdown or perform a balked an engine failure after the LDP, an approach to touchdown is dictated. Prior b or at LDP Balked Landing Go-Amtmd 1. Insrease collective pitch to adjust engine to 2. Cminute power (826" Tg and/or 111% torque). Maintain Nr within limits. 2. Accelerate to takeoff safety speed (V-) (52K3AS).

-

3. Landing gear up when positive rate of climb is established.

4. Adjust Nr to 96%with N2 beeper. Keep airspeed between V e s and best ate of climb speed, wbile oper;l.ting at 96% Nr. 5. When obstmction d e a r a ~ c is e asswed, accelerate to best rate of climb speed and continzre climb at 96%q. After LDP 1. hmease collective as ree@red (do not go beyand 2.50power) to cantain descent at9 &/or below 750 fpm.

engine

2. AtaboPt50lieetabavetoucbdrrwnbifiaieatpclicflare toabout 20" nose up.

3. Lwel nose to ' 5 to 10' nose up at dboat 30-feet above touchdown. 4. Apply collective to &hion hading.

5. After tooehdmm, necyclic and simPZtaneatrslp reduce C ~ Y ~ C tive to mimnm~Apply brakes as required

FAA APPROVED NO-ER 21,1978 Revised 3uly 26, 1982

K

FUGHT MANUAL

Part 1 Section III Emergency Procedures

SINGLE-ENGINE LANDING - CATEGORY "B" 1. Establish normal approach so as to arrive at 100-feet above touch-

down point at 50 K I M at a rate of no more than 500 fpm. 2. Decelerate to 50 feet a t 40 KIAS. 3. Continue deceleration to running touchdown at o r above translational lift. Use collective pitch to cushion touchdown. 4. After touchdown, neutralize cyclic and simultaneously reduce C O U ~ C -

tive to minimum. 5. Apply brakes a s necessary to stop within heliport confiies.

DUAL- E N G N FAILURE WARNING Rotor RPM will reduce t o an unrecoverable state with resultant loss of control unless autorotation is entered immediately.

Dual-enme failure requires immediate action for a power-off landing. The a varied conditions under which engine failure may occur prevent dictatiq standard procedure. However, a thorough knowledge of the helicopter's characteristics and emergency procedures will enable a pilot to respond correctly to the emergency. The altitude and airspeed at which engine failure occurs will dictate the action to be taken to effect a safe W i g . Should dual-en,@ne failure occur, a safe autorotative landing can usually be made. Upon dual-engine failure, the helicopter will swing to the left, due to the reduction in torque as e n m e power decreases. An immediate collective pitch reduction will be required to maintain Nr within safe limits. Minimum collective (full down) should be selected until Nr builds to within normal operation range. Apply tail rotor pedal a s necessary to maintain heading. DUAL-ENGM FAILURE WHILE HOVERING OR ON TAKEOFF AT 15 FEET OR BELOW Settling will be very rapid; however, the landing can be cushioned by increasing collective pitch a s the helicopter settles to the ground.

FAA APPROVED NOVEMBER 21, 1978 Revised July 26, 1982

~IKH?@~~~"X&

Pan 1, Section III Emergency Procedures

SA 4047-76- 1

FLIGHT MANUAL

Decreasins collective could cause an excessive sink rate. The helicopter should be held in a level attitude until is made with the gaud, then the cyclic conr stick should be moved slishtly forward of neutral. After ground is mad reduce collective pitch to minimum and apply wheel brakes. WARNING Excessive displacement of the cyclic control stick aft of neutral will decrease the main rotor blade-tail cone clearance and increase the possibility of striking the tail cone with a main rotor blade. DUAL-ENGINE FAILLRE DURING TAKEOFF AND I N m A L CLIMBAfter climb has been started. do this:

I

1. Immediately decrease collective pitch to minimum and establish a glide at 75 KIAS, if altitude permits. Regulate collective pitch to maintain rotor speed within limits. 1008 to lCS%,Nrshould be the optimum for glide to touchdown. 2. Cabin occupants - Alerted.

3. Landing gear - DOWN. UP over water if flotation system is not installed. 4. Make an autorotative landing following procedure in para-gaph titlecl

AUTOROTATIVE LANDINGS in this section. DUAL-ENGINE FALLURE DURING CRUISE A safe autorotative landing can be made if the helicopter is being flown at a safe altitude-airspeed combination and there is enough terrain clearance to reach a suitable landing area. When altitude and conditions permit, an air restart should be attempted. If neither engine will start, a normal power-off landing should be made as described in the para-gaph titled AUTOROTATIVE LANDINGS in this section. 1. Reduce collective pitch immediately and establish an autorotative glide at 75 KIAS. Regulate collective pitch to keep rotor speed within placard limits. A comfortable rate of descent will be reached if Nr is adjusted to 1008 - 105%.

2. Landing gear - DOWK. UP over water if flotation system is not installed. 3. Fuel and engine levers - OFF.

-

4. Cabin occupants Alerted

FAA APPROVED .NOVEMBER 21,1978 Revised May 15, 1992

6

-

FIJGHT MANUAL

Pan 1 Section III Emergency Procedures

5. If time and altitude permil, try to restan one or both engines. NOTE It will be left to the pilot's judgment. whether to attempt a stan or to make an autorotative ianding. If an attempt to stan is to be made, rtifer to the parabaph title h g i n e Restart in Right. ALTOROTATIVE LANDINGS Power-off autororative landing may be safely made except when operating at low altitude and at a low airspeed. These procedures apply after autorotative entry has been completed: 1. Esrabiish a glide at a minimum 75 KIAS.

-

2. Landing gear DOWK. U P over water if flotation system is not insralled.

3. At about 50 to 75 feet, execute a panial flare to about 20" nose up by moving ~yciicstick back with no change in collective pitch. This reduces airspeed and rate of descent and will cause an increase in rotor rpm. 4. Gradually increase collective pitch and apply f 0 r ~ d I dcydic as the helicopter

setties and until ground is made. 5. Maximum desired ground speed is 30 knots. Desired nose attitude is + 10"or iess. 6. Immediately after &round , decrease colieaive pitch smoothly and

siowly.

7. Apply wheel brakes.

FAA APPROVED NOVEMBER 21,1978 Revised March 2,1989

ENGINE COh'TROL MALFUNCTIONS

N2 COmROL ACTUATOR RUNAWAY symptoms:

I'niLial:

Large torque spli~

Confirming: Both engines above idle. Malfunctioning engine N2 a 97-99911 (iowside failure) or 107% (highside failure). Malfunctioning engine does not respond to N2 beep commands. Both torques respond to collective motions.

Action:

Beep good engine N2 to match torques.

NOTE Dual-engine operation below 100% N2 is authorized without airspeed restriction when executing the emergency procedure for an a c t d iowside failure.

This maifunction does not cause any lass of available engine power.

FAA APPROVED NOVEMBER 2 1.1 978 Revised March 2,1989

I

Part 1Section EX -erg, mc-es

AFUGKT i MANUAL @ ~ ~ 5 ~ ~ ~ f

1 COLLECTIVE BIAS ACTUATOR OR DROOP COMPENSATOR AMPLIFTER ~ D C A )FAILURE

I~~itial:

Large torque split.

Confirming:

Both engines above idle. Both engines respond to Nq beep commands.

does not respond or responds One engine tosluggishly to collective motion. Nr responds to collective motion. Action:

Match torques with X2 beeper. Beep to compensate Ng and torque match with collective motion. Avoid large w i d power applications.

ENGINE COF??OLFAlLS TO BIGH POWER Ehgine control failure to high power wi3.l cause a large torque split and could be misinterpreted as an engine power failure. It must be noted tbat in a control failure to high power the Nr would increase rapidly, indicating that power has been added to the system. Symptoms:

Initial: Conf'iing:

Large torqae split. Aigh torque engine at o r above topping N1. Nr increasing

Action:

- possibly above limits.

On the gromtd, pull both engine levers to shut-off.

In hover, the helicopter will rise in altitude as Nr increases. Pull back the engine lever on the affected engine as necessary to appro+tely match oatpat torques and regain control of Nr. Proceed as d+ scribed below to control the Nr and torque during

l=ws

In fligtrt, increase collective pitch as necessary to conbin Nr within limits- When Nr is stabilized, tetad the enpine l w e r of the affected engine as necessary to reduce output toque to approximately 15% torque less than the good engine (thiswill allow same margin for cwrdinaling collective pitch and engine l w e r movements). Use the following procedure to help &maintain Nr w i t W l mts and land as practicable.

FAA APPRONOVEMBER 21, 1978 Revised July 26, 1982

FUGHT MANUAL

w

-

Collective pitch changes should be made in steps so that the output torque of the affected engine is always less than, or equal to, the torque of the good engine. For climbs, increase the collective pitch first, then increase the engine lever. For descents, decrease the engine lever first, then decrease the collective pitch. Nr should be kept within limits throughout. Depending upon the workload involved in maintaining Nr within limits using the above procedures, the pilot might consider shutting down the affected engine as means of reducing the workload and containing the emergency situation. See Engine Shutdown in Flight in this section. Avoid rapid collective reduction at all times. If the malfunctioning engine is not shut down before landing, the pilot should be sure that upon landing, the collective is not fully reduced before shutting down the engine.

The greatest hazard of this malfunction is an overspeed of the engine, transmission, and rotor systems. This overspeed condition will be prevented by the mechanical overspeed system at 118% '02. If the overspeed system is tripped by a rise in N2, a sharp cutback in all parameters of the affected ensine will be noted. Cycling at the overspeed trip point may be noticed. The pilots should not depend upon the overspeed system for containment of high Nr.

ENGINE AUXILIARY SYSTEMS FAILURE ENGLNE OIL SYSTEM FAILURE Symptoms:

#1 or #2 ENG OIL PRESS warning light on and engine oil pressure below 50 psi, or engine oil temperature above 107°C.

Action:

Transition into single-engine flight if possible. Shut down affected engine and land as soon as practicable. If single-engine flight is not possible, use N? beeper to reduce output power of affected engine to minimum required. Land as soon as possible.

FAA APPROVED NOVEMBER 21,1978 Revised May 15, 1992

Part 1Section III Emergency P T O C ~ ~ U ~ ~ S FUGHT MANUAL

AIEHRHM~~~ NOTE

A single indication of low oil pressure should not be treated as an oil system failure unless i t can be verified that the second indicating system has failed or there is some other failure indication, EL'EL PRESSURE WARNKG LIGHT ON Dlumination of the FUEL PRESS warning li@t indicates that the engine fuel pressure at the h . 3 pressure pump inlet has fallen below 5 psi. The fuel pressure gauge (if installed) provides a secondary indication of engine fuel pressure and should be cross checked to confirm actual fuel pressure. ,e Loss of fuel pressure could be caused by blockage or failure of the en* low pressure ejector pump. The engine should operate normally with an ejector pump failure, but the condition must be corrected prior to the next flight. Low fuel pressure fluctuation might be caused by air in the engine fuel line, which could be caused by a leak in a fuel line or impending fuel starvation. Such causes will usually result in unsteady engine operation, including fluctuation in output torque, XI. Tg and possibly unsteady rumbling noises. Symptoms:

-

Initial: 21 ENG o r +2 ENG FUEL PRESS warning light OK

Confirming: Fuel Pressure gauge below 5 psi AU other engine parameters normal Action:

Note discrepancy.

Correct prior to next flight.

If other engine parameters indicate a possible en*,e

problem, adjust flight conditions to minimize impact of possible engine failure. Land as soon as practicable. E'LTEL FLOWMETER BEYObD LIMITS (IF FUEL FLOWMETERS ARE INSTALLED If a fuel flowmeter indicates erratically or goes beyond limits, crosscheck the other engine indicators (XI, T5, torque, fuel pressure caution light). If all other indications a r e normal, continue the flight and monitor the other engine instruments closely.

I EA?S

CAUTION I3GHT ON

The EAPS caution light goes on in the following cases: (A) EAPS switch OX and EAPS doors not closed or EAPS blower bladed or EAPS pressure swktch activated.

(B) EAPS switch OFF and EAPS doors not open.

FAA APPROVED XOVEMBER 21, 1978 Revised AUGUST 27, 1975

-

C - g: Actian:

EAPS advisory light off. Reposition aircraft to environment which does not require EAPS, trim EAPS OFF if desired, land as socm as practicable.

s'ymptoms:

EAPScaation&hton, EAPSsraitchOFF.

Action:

Suspect me or both EAPS doors not fully opened. Note T5 Herences for matched torques, pfan on

reduced power available an affected engine. ANTI-ICE CAUTION LIGHT ON The anti-ice cauticm light for each engine is on in the fallowing cases:

A) Anti-ice switch ON and EAPS doors OPEN or engine anti-ice bleed-air failed

B) Anti-ice switch OFF.and engine anti-ice bleed air ON. Symptoms:

Either engine anti-ice cadion light ON-tbat engine anti-ice sw&ch ON tbat engine anti-ice advisory light O F F

Actian:

Assumeanti-iceWhreonihaiagin+land as soon as padieable

anticipate possible ice ingestian (if in inadvertent icing candftim) on affected engine. Symptoms:

Either engine aPti-ice cautim ON-&at engine anti-ice advisory light ON-that engine anti-ice swikh OFF,

A ~ ~ ~ c A L . * Assmne anti-ice bleed air is flawing on afEected engine. Check AI CONT circuit breaker in and cycle --ice switchto~mpttoresetazlti-ice?ave*'EresetisrmwecessEuI, plan an s-y &ed power available m aflected el@ne.

(Helicopters eqrripped with mow pmtecliallBit) The engine anti-ice caution light d tke affected engine willgoanifthehlethgaterpadsoftheicepmteciian kit should overheat o r fail to achieve proper opetemperatrve witen the respective enghaanti-ice switch is an. The engine --ice advisory light will remain an when this occurs indicating proper opemon of the enanti-ice bleed air system.

me

FAA APPROVED NOVEMBER 21, 1978 Revised MAY 22, 1980

3-6C

J

FUGHT MANUAL

Part 1 Section III Emergency Procedures

ENGINE CKIP DETECTOR CAUTION LIGHT ON

Symptoms:

=IENG CHIP o r =2 ENG CHIP caution light on.

Action:

Pull and reset CHIP DET circuit breaker of affected engine.

E light remains OFF, continue flight. If light comes back on reduce the affected engine output

power to the minimum required for flight by using the N2 beeper and land as soon a s practicable.

If light is accompanied by abnormal noises, oil pressure, or temperature, and single-engine flight can be maintained, shut down affected engine and land as soon as pracucable, If single-engine flight is not possible, reduce power of affected engine to the minimum and laad as soon as possible. MAIN GEAR BOX SYSTEM FAILURES

The most common main gear box system makfmxtions a r e of three general types. (1) lubrication system failure, (2) transmission component failure and (3) transmission accessory failure (to include the transmission oil cooler). The pilot has a cockpit gauge indicating oil temperature and pressure and three caution lights indicating low oil pressure, high oil temperature and chip detected, which inform him of the operating condition of the system. A thorough knowledge of the main transmission and accessory systems w f l l enable the pilot to epaluate the indications available to him, diagnose problems and act accordingly. In generdl, a single failure indication dictates that the helicopter be landed as soon as practicable. If multiple failure indications, including abnormal noise and/or vibration, are noted, land immediately.

-

GEAR BOX OIL SYSTEM FAILURES

Main Gear Box Oil fump Failure

Spmptoms:

Oil pressure falls to 30-33 psi and remains steady.

Action:

Land as soon as practicable. Monitor main transmission gauges aad caution lights for further failure inbications.

NOTE The remaining pump w i l l normally provide adequate lubrication for the main gear box

FAA APPROVED NOVEMBER 21, 1978 Revised FEBRVARY 22. 1985

'

Part 1, Section III

&/K@BSKM~&

Emergency Procedures

FUGHT MANUAL

Low Oil Pressure Symptoms: initial:

MAIN XMSN OIL PRESS warning light - on

C o n f i g : Main transmission oil pressure less than 20 psi. Actions:

Over Land Land as soon as possible. Descend to a minimum safe altitude and reduce airspeed to 80 knots if required to reach a landing site. if loud or unusual noises or hi$ vibration levels occur, land immediately. Over Water Descend to a minimum safe altitude, proceed to the nearest landing site, and reduce airspeed to 80 knots. Land as soon as possible-

If loud or unusual noises or high vibration levels occur, land immediately. Most likely cause of low oil pressure is an excessive loss of nans- ( mission oil. Other indications such as high temperature or chips detected may or may not occur. Do not anempt M e r flight until cause of low pressure indications has been corrected. Hizh Oil Temperature

initial:

,

MAIN XMSN OIL HOT warnin2 light - on.

Confirming: Main uausmission oil temperature above 105°C. Actions:

Reduce power, and land as soon as practicable. If oil temperature rises above 120°C or if oil pressure drops, land as soon as possible. NOTE If the indicated oil temperature remains above 105°Ca transmission oil cooler malfunction is indicated. Transmission oil temperature should be controlled by avoiding hish power settings andor prolonged hovering. FAA APPROVED NOVEMBER 21,1978 Revised June 17, 1997

Main Gear Box Chip Detector The main transmission is equipped with "Fuzz Burn" type chip detectors. Symptom:

MAIN XMSN CHIP caution light - on.

Action:

Pull and reset MGB CHIP DET circuit breaker. If light remains off, continue flight If light comes back on, land as soon as practicable.

In case of secondary failure indications, such as an oil pressure loss or temperature rise, land as soon as possible. Operate at best rate of climb speed (minimum power required) and at a safe minimum altitude. If loud noises or high vibration levels occur, land immediately.

IN'TERMEDJATEOR TAIL GEAR BOX CHIPIHIGH OIL TEMPERATURE DETECTED Intermediate and tail gear boxes have combination high-temperature and fuzz-burn chip detecton. Symptom:

INT GB CHIPlHOT or TAIL GB CHlPMOT caution light on-

Action:

Pull and reset the CHlP DET I/TGB circuit breaker. If light remains off, continue flight. If light comes back on, land as soon as possible. If loud noises, smoke or high vibration levels occur, land immediately. Anticipate tail rot& loss of thrust. (See TAU. - ROTOR MALFUNrnONS .)

FAA APPROVED NOVEMBER 2 1, 1978 Revised March 16, 1994

FUGHT MANUAL

V

.

ROTOR BRAKE CAUTIOK LIGHT ON - ROTOR T U N I N G The ROTOR BRAKE caution light indicates that pressure has been applied to the rotor brake system. If this pressure results in a rotor brake puck dragging against brake disk, an overheated brake disk and possible rotor brake fire might occur. If the ROTOR BRAKE caution light is on and the rotor head continues to turn, do this: Symptom:

ROTOR BRAKE caution light on Rotor head m i n g

Action: During start or shutdown on gound Shut down engine(s) or release rotor brake During Flight Rotor brake switch - Release - hold until caution light goes out. or 5 seconds. For manual rotor brake installations - check handle in detent. If ROTOR BRAKE caution light remains on, investigate for rotor brake fire (smoke, unusual noises or vibrations etc.). If rotor brake fire is suspected, land immediately. FLIGHT CONTROL HYDRAULIC SERVO SYSTEM FAILURE Complete control of the helicopter can be maintained through either stage of the flight control system; however, prolonged operation with only one servo system operating is not recommended. This is an emerzency condition, since conmol is not possible with both servo systems inoperative. Fli$t conditions and route should be chosen which will allow a safe immediate landing. PUMP FAILURE OR LOSS OF PRESSURE IN BASIC HYDRAULIC SYSTEM Symptom:

# 1 or #2 SERVO SYSTEM caution light on (below 1600 psi).

FAA APPROVED NOVEMBER 2 1, 1978 Revised May 15, 1992

Part 1 Section III Emergency Procedures

FLIGHT MANUAL

Confirming: Decrease in corresponding gage pressure. Action: 1. Servo switch Failed system OFF if no usable pressure remains o r if erratic pressure is observed. 2. Reduce airspeed to 125 KIAS o r lower.

-

3.

Land as soon as possible. 4. If controls bind o r a r e restricted'. land immediately. XOTE If the second see has malfunctioned and flight range considerations permit. immetiiately lower the landing gear. If gage pressure has decreased to below about 2000 psi. i t may be necessary to use the emergency landing gear system to lower the landing gear. The variable vibration absorber may also be inoperative. SERVO UKTT JAM OR MALFUNCTIONING SERVO SHUTOFF VALVE

I

NOTE For helicopters equipped with a tail rotor servo shutoff valve PK 76076-65003-011 refer to Supplement No. 28 for proper procedures. Symptom:

+ 1 or 8 2 SERVO SYSTEM caution light on and corresponding hydraulic =e

pressure normal

-

pressures normal Confirming: Both system Both flight control servo switches ON Action: 1.

-

2.

Reduce airspeed to 125 KlAS o r lower. If binding o r restrictions in the controls a r e felt, shut failed system and land as soon as possible.

3.

If binding or restriction in the controls does not disappear after failed system is shut off. land immediately.

4.

aff

If no binding o r restriction in controls is felt, recycle appropriate SVO JAM circuit breaker on DC ESNTL bus . If appropriate SERVO SYSTEM Light goes off, a momentary jam is indicated. Continue for a landing as soon as practicable.

5. If SERVO SYSTEM light remains on after step 4., either a jam o r malfunctioning servo shutoE valve is indicated. Pull appropriate SV0 SHUTOFF circuit breaker. This removes power to SERVO SHUTOFF VALVE and, if valve or shutoff control switch has rnaJfunctioned, hydraulic system w i l l be restored. 6. Cycle SV0 JAM circuit breaker to reset holding circuit on warning light. 7.

If SERVO SYSTEM light remains on after step 6 is completed,

..

continuing servo jam condition is indicated; reset SVO SHUT0 FF circuit breaker, and land a s soon as possible. FAA APPROVED NOVEMBER 21. 1978 Revised DECEMBER 12, 1986

SNK@BS~YIE Part 1Procedures Section 111 Emergency FUGHT MANUAL

8. If light goes off after step 6., a malfunctioning servo shutoff valve has been isolated. Land as soon as practicable. NOTE When a SERVO SYSTEM caution light goes on in flight, a holding relay will keep the light on even if the cause was momentary. Since the holding relay operates through the landing gear interlock that functions only when the helicopter is airborne, the caution light will go off upon landing if caused by a momentary jam condition. The caution light may be cleared in flight by recycling the appropriate SVO JAM circuit breaker on the DC ESNTL bus . FLIGHT CONTROL DAMPER JAM Symptom Cyclic or collective control will not move with normal hand pressure. Confirming: Typically, only one axis is involved and remaining axes are normal. Action: If on the ground, shut down and investigate. In flight: 1. Check that cockpit control travel is not impeded by a foreign object. 2. Ifjam persists, turn off appropriate trim system. 3. If system is still jammed, apply increasing pressure to the affected con-

trol to sever the shear device in the control system damper. This action should be taken against trim to possibly reduce overshoot. Typical amount of force required is: Collective Pitch Roll

55-62 pounds 17-20 pounds 35-40 pounds

PITCH BIAS ACTUATOR RUNAWAY Actuator runaway is characterized by a slow repositioning of the actuator to its stops accompanied by a corresponding change in pitch attitude. Confirmation is indicated by observing the indicator pointers on the AFCS channel monitor with the channel monitor switch in PBA. At above 45 knots, the pointer will be at the full left or right extremes, indicating that the actuator has moved to its extended or retracted stop. Below 45 knots, the system is not operational and the pointers should be full left. Recovery is made by repositioning the cyclic stick. Electrical power should be removed by pulling the BIAS ACTR circuit breaker.

FAA APPROVED TU'OVEMBER 21,1978 Revised September 7,1990

3- 10A

I

Part 1 Section 111 Emergency Procedures

&FUGHT % i BMANUAL ?~~fl& NOTE

Pulling BIAS ACTR circuit breaker will cause loss of airspeed hold with AFCS Phase 111 and loss of Droop Compensator Amplifier (DCA) system redundancy in all configurations. AFCS (PHASE 11) FAILURE In IMC, if any AFCS malfunction results in less than one pitch and one roll channel remaining functional, maintain airspeed below 120 KIAS. Attempt to transition to VMC. If at least one channel of yaw AFCS is not functional, reduce airspeed as necessary. Failures that may occur include: single-channel hardover, single-channel actuator failure, or oscillatory malfunction. SINGLE-ACTUATOR HARDOVER A single-actuator hardover in the pitch and roll axis will result in an abrupt change in attitude. The attitude change in yaw is less abrupt and may not be detected as easily by the pilot. The failed axis may be determined by turning the channel, monitor switch on the AFCS indicator to the failed axis position (PITCH, ROLL, or YAW) and monitoring each actuator indicator for an off-center indication. Recovery is made by repositioning the controls to the trim attitude and turning the failed axis off on the appropriate AFCS control . SINGLE-ACTUATOR FAILURE Failure of an actuator to reposition, in any axis, is indicated by failure of the indicator needle of the corresponding channel to move. A single-actuator failure may be determined by turning the channel monitor switch to the PITCH, ROLL, or YAW positions and monitoring each actuator indicator for an off-center indication. The failed axis should be turned off on the appropriate AFCS control . OSCILLATORY MALFUNCTIONS Cyclic stick jitter or lower frequency lateral or longitudinal stick oscillation, -ztbether or not accompanied by corresponding aircraft response, may indicate a malfunction of a vertical gyro or an electrical power problem. Disengage individual SAS channels until an improvement is realized, commencing with the axis displaying the most stick or aircraft motion. Subsequently, reengage those channels not contributing to the problem. AFCS (PHASE 111) FAILURE In IMC, if AFCS A/S advisory light is not on at airspeeds above 50 KIAS, move cyclic trim switch from AUTO to ON and use AFCS Phase I1 minimum IMC airspeed of 60 KIAS and recommended approach speed of 70 to 125 KIAS.

FAA APPROVED NOVEMBER 21,1978 Revised March 2,1989

OUTER LOOP SRUTDoWhT Symptom: . A R O SRUT-DN caution light on. Action:

Turn the channel monitor switch on the AFCS indicator throw the PITCH and ROLL positions, obsemhg at each position any mismatch between No. 1and No. 2 channels. If a mismatch is observed select cydic trim to OX. Sltbsequently turn off the affected inner loop channel. If no inner loop discrepancy was observed, do the following sequentially until caution light goes off and system is restored to operation. (See Figure 3-1-01

-

-

Yaw trim switch OFF. If AUTO SHUT DN light goes off. the shutdown was yav isolated. Retarn yaw trim switch to 0% and contirme flight. If light did not go off, place cyclic trim switch ON. If AUTO SHUT DX light goes off, the shutdown was pitch axis related. Return,cyclic trim switch to AUTO and continue

-

nmt. If the light did not go off in either of the above checks, the shutdown was roll axis related, m w e cyclic trim switch to AUTO and place y a w trim switch OFF far 1 second and retrvnto ON and cmtinue £light. If the shutdawn feature subsequently causes repeated emmcia-

tions, the pilot may choose to degrade to a partial AFCS

FAA APPROVED NOVEMBER 21, 1978 Revised September 22, 1982

Part 1 Section m Emergency Procedures

&K@RHKOY/&.

FUGHT MANUAL

-7 PLACE YAW TRIM MASTER SWITCH OFF

YES

AUTO SHUT-ON LIGHT ON?

YAW CHANNEL

N PUCE YAW TRIM lllASTER SWITCH ON

I

PLACE CYC TRIM MA!ZER m C H FROM AUTO TO ON

1

ROLL CHANNEL

F i e 3-1-0 AFCS Ouker Loop Shutdown

Phase III system using the above traubleshooting sequence. For -le, if the pitch axis is suspected, the pilot may select AFCS Pbase II in pitch and roll by placing the cyclic trim switch ON and cantinue flight using the trvn coordination or heahold portian of AFCS Phase III. FAA APPROVED NOVEM%ER 21, 1978 Revised September 22, 1982

~FLIGHT E MANUAL ~ ~ Emergency I PartE1,Procedures Section BIII SINGLE AXIS OUTER LOOP HARDOVER The purpose of the auto shutdown portion of AFCS Phase I11 is to disable the outer loop before a hardover can develop. Without auto shutdown, an outer loop hardover would cause the affected control to move in the direction of the hardover, resulting in a corresponding pitch, roll, or yaw excursion. Recommended recovery is made by repositioning the controls to establish trim attitude, and turning off AFCS Phase I11 cyclic or yaw as appropriate. If the airspeed hold function is on, an airspeed transducer malfunction may result in a pitch attitude excursion similar to that expected with a Pitch Bias Actuator runaway. Recovery is made by repositioning the cyclic stick for desired pitch attitude. Subsequently, turn cyclic trim to ON and check PBA indicator for position appropriate for the aircraft trim condition. If indicator position is abnormal, refer to PITCH BIAS ACTUATOR RUNAWAY in this section. An altitude loss of 200 feet was demonstrated during recovery from a forward pitch AFCS Phase III hardover initiated at 155 KZAS level flight. TAIL ROTOR MALFUNCTIONS Tail rotor malfunctions can be classified generally as drive system or control system failures. The helicopter's reactions and consequences of tail rotor failures will vary widely, depending upon the type of failure and the flight condition of the helicopter at the time of the failure. In general, however, control system failure is less critical than drive system failure. TAIL ROTOR DRIVE SYSTEM FAILURE IN FORWARD FLIGHT Since tail rotor drive system failures may produce a large response from the helicopter it is very important to recognize impending failure. The failure will usually be preceded by excessive noise or vibration in the tail section. When this occurs at high speed, immediately slow to 100-1 10 K I M . Tail rotor drive system failure is always accompanied by a sharp yaw to the right. Reduce collective as necessary to control the yaw. Autorotate if necessary to control yawing rate. When a landing is to be made: 1. Autorotate 2. Maintain an airspeed of 75 KZAS 3. Landing gear - DOWN. Up over water if flotation system is not installed. 4. Engines - Shutdown before touchdown

5. Execute autorotative landing NOTE Ground speed must be held to a minimum to reduce the tendency to roll over due to yaw. FAA APPROVED NOVEMBER 2 1,1978 Revised October 19,2001

1

Part 1, Section III Emergency Procedures

& #FLIGHT f m ?MANUAL B??~fl~@

TAIL ROTOR DRIVE SYSTEM FAILURE IN A HOVER 1. Maintain helicopter attitude and attempt to achieve zero groundspeed and drift. 2. Engine levers - OFF at about 10 feet if possible. 3. Increase collective to cushion landing.

TAIL ROTOR CONTROL SYSTEM FAILURE

I

The most probable tail rotor control failures are jammed or binding controls, before or after the mixing unit. An interruption of the control system such as a severed control cable is also possible. The helicopter's reaction and the cues available to the pilot will vary widely with flight regime. Fixed Pitch Setting - General Depending upon the origin of the interference, the pedals may not be movable (control bind or interference between the pedals and the collective to yaw control mixer), or movable through a restricted range (blocked pedal damper orifice). Full collective control is available under these conditions and normal collective to yaw inputs to. tail rotor pitch will be available. A control bind or interference aft of the mixer output will result in a fixed tail rotor pitch situation. Full collective control is available when this occurs only if the pilot allows the pedals to move simultaneously with collective. No tail rotor pitch control is available, and adverse yaw proportional to the amount of collective change should be expected. Therefore, in this fixed tail rotor pitch condition, disregard pedal motion as an indicator of potential tail rotor control. The technique for landing the helicopter will vary greatly, depending upon the resulting tail rotor thrust. A practice approach at altitude may be useful in predicting helicopter yaw response as a function of airspeed, but care should be taken to avoid slowing beyond safe limits. Yaw SAS will continue to provide short term damping reducing gust related disturbances and should be selected on if so equipped. In general, touchdown will be made with some yaw. Attempt to land with a minimum rateof-descent. Touchdown should occur first on the most forward main landing gear. Counter roll tendency with cyclic while allowing nose of helicopter to pivot toward direction of landing. After touchdown, slowly lower collective to minimum and use differential braking to control helicopter on ground.

I

Fixed Pitch Setting - Reduced Power The helicopter will swing right as power is applied, with the amount and rate of yaw proportional to the amount and rate of power applied. Decrease power as necessary to control the heading of the helicopter. Make a roll-on landing at the speed, up to 40 knots, which results in minimum yaw on touchdown.

3- 12

FAA APPROVED NOVEMBER 2 1,1978 Revised JUNE 29, 1984

-

Fixed Pitch at High Power Settings The helicopter will left as collective pitch is reduced, at a rate and amount proportional to the rate and amount of collective reduction. Make an approach and landing, accepiing some left yaw on approach. Touch down at a speed slow enough to permit an increase of collective at touchdown to minimize yaw. Consideration might be given to selecting a runway with a right crosswind which will reduce some of the adverse yaw on touchdown. Fixed Pitch Sette

at Cruise Power

The helicopter will left or right a s power is reduced or added, although in general the reaction should be milder than in the previous cases described. Plan approach and landing to touch down w i t h power a s close a s possible to cruise set-. Tail Rotor Cable Failure

- Centering Spring Installed

Upon recognizing a tail rotor cable failure, proceed to a landing area offering at least 2,000 feet of surface. Crosswind landmgs with winds greater than 10 knots are not recommended. Yaw SAS continues t o provide short term damp* that reduces gust related disturbances and should be selected on. If insufEicient yaw control remains to enable trimmed flight, accept the sideslip and fly the helicopter along a desired track usirig laterial cyclic. A single cable failure will affect tail rotor control from the centering spring position (-2O) in the failed direction only, leaving f u l l control in the opposite direction by means of the remaining cable. Although the helicopter may continue to fly normally enroute and during descents, it is recommended that the following ciual cable failure landing procedure be followed to complete a roll-on lanas soon as the hedcopter heading is satisfactorily aligned with the runway.

Fixed Pitch Setting

- Tail Rotor Centering Spring (-2')

Initiate a long shallow approach, commencing descent a t about 80 KIAS. Plan approach to be established low, on short final, with 60 KIAS. Expect heading to remain left of track but responsive to collective changes; yawing further left with collective reduction and right with collective increase. When over the landing surface, arrest remaining rate of descent with a gradual cyclic flare that results in a level flight deceleration to 40 KIASUse small collective inputs as necessary and continue a level flight deceleration at 10-feet wheel height or lower. Left yaw will decrease, as airspeed decreases and power required increases. Plan touchdown point to occur as helicopter heading becomes ali,gned with runway. Do not atkempt to slow beyond where an acceptable yaw attitude is realized. After touchdown, gradually reduce collective while maintaining runway he* with Werentia braking.

FAA APPROVED NOVEMBER 21, 1978 Revised JL.. 29, 1984

I

Part 1 Section 111 Emergency Procedures

FUGHT MANUAL

A balked landing is executed during the approach by adding sflicient collective to initiate a gradual acceleration while establishing a shallow climb. As airspeef increases, additional collective may be used if a steeper climb is desired. A balkec, landing should not be attempted once the nose swings right of center. Collective application beyond this point can result m a continuing flat rotation to the right. CAUTION Large control inputs at moderate to rapid rates are to be avoided. Collective applications, especially, should be slow and deliberate, not exceeding those required to gradually transition from one flight regime to the next (approach to balked landing, descent to level off, and touchdown to rollout are examples).

FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

Part 1 Section 111 S~%!@RHKOH& E~~~~~~~ procedures FUGHT MANUAL

TAIL ROTOR DAMAGE Damage to the tail rotor can be noted by a sharp increase in tail rotor vibration. Slow to about 90 KIAS and land as soon as possible. FIRE ENGINE COMPARTMEKT FIRE Symptom:

I

Initial: FIRE warning light and T-handle light on with continuous tone heard. NOTE Helicopters modified by PK 7607O-j5049-Oll are equipped with individual #1 ENG FIRE and #2 ENG FIRE master warning lights. Confirming: Smoke, burning odor, or erratic engine indications. Action: Reset tone by pressing a FIRE warning light.

.

1

With fire confirmed, establish safe single-engine flight while fully retarding illuminated T-handle of affected engine.

I

Place fire extinguisher switch to MAW. An extinguishing of the appropriate FIRE warning light and T-handle is ive information that the fire has been arrested. If fire persists, place fire extinguisher switch to RESERVE.

I

If any sign of fire still persists, land immediately otherwise, land as soon as possible. Shut down and leave helicopter immediately. Internal Engine Fire After Shutdown Symptom: Initial:

Smoke, flames, andlor a rapid increase in Tj.

Confirming: No FIRE warning light or tone. FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

Actior :

With engine lever OFF, press starter button to motor engine. If fire persists, fully retard T-handle and continue to motor en,$ne while f i r e guard discharges an extinguisher into engine a i r intake. If fire still persists, shut off all switches and leave helicopter.

NOTE If fire warning lights should go on at any time during this procedure, follow procedure in the paragraph titled ENGINE COMPARTMENT E m .

Cabin o r Cockpit Eire 1. Begin a descent, if able. 2. Close all cabin ventilation.

CAUTION If smoke accumulation becomes a problem, execute

Smoke and Fume Elimination procedure below and continue to fight fire. 3. Heater BLEED AIR switch o r ECU CONT SEL switch

- OFF.

4. Determine location and extent of fire. Use portable fire extinguisher.

5. Land as soon as possible and shut down.

6. Iffire is not out, leave helicopter.

Bagawe Compartment Smoke Detected

)

Synptorn:

-

SMOKE DET BAGGAGE warnkg

light illuminated.

Confirming: finusual smell of smoke. Action:

compartment Land as soon as possible. Inspect -gage and fight fire with fire e x t i n , g h e r if required. NOTE The b a g m e compartment is sealed so that harmful quantities of smoke will not enter the cockpit and cabin. Large quantities of smoke in the bagcompartment will probably be smelled in the cockpit/cabin area.

FAA APPROVED NOVEMBER 21, 1978 Revised OCTOBER 14, 1980

FLIGHT MANUAL

-

-

Smoke and Fume Elimination The following procedures will eliminate smoke and fumes from the cockpit and cabin.

-

1. Pilot and copilot overhead vents OPEN.

2. Pilot's clear view window - OPEN.

3. Heater BLOWER switch or ECU CONT SEL switch - ON. Turn OFF if smoke or fumes is noted coming from cabin or cockpit air outlets. Elecmcal Fire Elecm~alfires are generally indicated by a smell of burning insulation andlor acrid smoke. As in any fire, the most important consideration is to maintain safe flight while dealing with the fire. When detecting an electrical f ~ ethe , pilot must shut down any unnecessaq electrical equipment.

If in VMC, perform the followin,o: Turn off both dc generators. the ac generator, and the inverterfs), and pull dc bus-tie circuit breakers (2) on the dc primary circuit breaker . Choose the appropriate action. Action:

- If fire persists, turn on all generators and inverter(s), and nun off battery. The fire should go out. Land as soon as possible. The emergency blowdown system will be required for landing gear extension.

-

If fire goes out, restore the following equipment in sequence. Allow a brief interval benveen actions to assess response. Disengage the component if signs of fire return. Land as soon as possible.

- AC generator - NO. 1 dc generator, and if fire remains out, reset the bus-tie circuit breakers (2)If in IMC flight, transition immediately to VMC flight if possible. If it is not possible to achieve VMC flight and the source of the fire is not evident, see Figure 3-1.

FAA APPROVED NOVEMBER 2 1, 1978 Revised May 15, 1992

/

Pan 1 Section 111

'fl'@~gfyff

FLIGHT MANUAL

Emergency Procedures

ELECTRICAL FIRE PROCEDURES. IMC

START HERE

e l AND k2 DC

PROBLEM IS I N THE nY2 GEN. LAND AS SOON AS PRACTICABLE.

FIRE

GENS OFF

I

FlRE

?/)///

PROBLEM IS I N THE $1GEN. LAND AS SOON AS PRACTICABLE

4 1 DC GEN OFF 6 2 DC GEN ON

LOSS OF ALL , 5 DC PWR. , IMMEDIATELY 'PROCEED TO ( NEXT BOX. , #

/

FlRE

/ /

I

I

PROBLEM IS I N THE n Y 1 OR nY2 DC PRIMARY BUS. LAND AS SOON AS

FIRE

PROBLEM IS I N THE BAT. LAND AS SOON AS PRACTICABLE

4 1 AND *2 DC

GENS ON. BAT SW OFF. I FIRE

I

BUS TIE CB'S PULL. PROBLEM IS I N THE ESSENTIAL BUSLAND AS SOON AS POSSIBLE.1.2

1

BUS TIE CB'S

E%E&

AC GEN OFF. ON.

-1 4

PROBLEM IS I N

liE E k:Et& AS PRACTICABLE.1

7 FIRE

PROBLEM IS I N A N AC BUS. LAND AS SOON AS POSSIBLE. 1,2

1

I 1. IF REQUIRED BY WEATHER CONDITIONS. BUS SYSTEMS MAY B E RESTORED SINGLY BY USE OF APPROPRIATE CIRBREAKERS ON BOTH AC PR~MARY &26v BUSES AT THE SAME TIME. 2. IF ABLE TO OBTAIN VMC. LAND AS SOON AS PRACTICABLE AFTER SECURING ELECTRICAL SYSTEM.

Figure 3- 1 FAA APPROVEDNOVEMBER 2 I, 1978 Revised March 2. 1989

Part 1, Section 111 Emergency Procedures

~K ZZRSK~~BE FUGHT MANUAL

LANDING GEAR MALFUNCTTONS Normal gear extension is indicated by the lighting of all three green gear dowi. indicator lights apd the red UNLKD gear warning light going off. Complete retraction will result in all four lights ,ooing off. Any other indications must be assumed unsafe and require the following emergency action. NOTE Once a safe down indication is obtained, the landing gear should not be cycled. Land gntly, have the gear pinned, and shut down. Taxi operations with unpinned gear should be avoided. Also avoid the use of the rotor brake with unpinned gear. If sear cannot be made to extend completely or cannot be pinned, a landing on a cradle or similar fixture should be made (mattresses, sandbags, etc.). Asymrnemcal gear extension could increase roll tendency. Landing with all wheels up would probably not lead to roll over. Consider evacuation of engers and unrequired crew while at a hover; avoid rapid deceleration of rotor. LANDING GEAR WILL NOT EXTEND NOTE Absence of normal No. 2 hydraulic pressure or dc essential bus power requires the use of the emerzency landing gear blowdown system. Symptom: On lowering the gear handle, none of the three s e e n indicator lights illuminate within the normal extension period. Action: Ress to test the cautiontadvisory and check landing gear lights. Slow to about 80 KIAS.

Check that LDG GR CONT and LDG GR POS LTS circuit breakers are set. Attempt to cycle the gear handle several times and leave down.

FAA APPROVED NOVEMBER 2 1, 1978 Revised May 15, 1992

Part 1 Section 111 FLIGHT MANUAL

On final approach before reaching a hover, check that the normal gear handle is selected Dh', and then turn and pull the emergency landing gear T-handle to discharge the blowdown bottle. A three green safe indication is to be expected following use of blowdown. NOTE Once blowdown bottle has been discharged, do not attempt to raise landing gear. Maintenance is required prior to next flight. Complete the landing by touching down lightly. Have a crewman pin the gear prior to taxi or shutdown. UNSAFE INDICATION - GEAR EXTENSION Symptom: On lowering gear handle, one or two of the three green indicator lights do not light, and/or red UNLKD light remains on. Action : If a green light is off, press test button on caution-advisory . If light does not test, switch light bulb with one of other lights and check for a safe indication. Slow to about 80 KIAS. Check that LDG GR C O W circuit breaker is set. Cycle landing gear several times. CAUTION If any indications of a jammed gear are present, leave gear handle DN. If landing gear still does not extend, hover and have someone outside examine landing gear. Consider having ground crew make a careful attempt to pull gear into place. Use of the emergency blowdown system is not recommended for cases with mixed safe and unsafe indications.

FAA APRROVED NOVEMBER 21,1978 Revised September 7,1990

Part 1 Section 111 Emergency Procedures

I

FUGHT MANUAL

If at this point the gear appears to be extended and/or a safegear indication i s obtained, touch down lightly and have crew pin gear. Do not put weight o r taxi on unpinned gear.

UNSAFE INDICATION - GEAR RETRACTION

Flight above 130 KIAS is not authorized with the landing gear extended or partially extended. Symptom:

I

With gear handle UP, &e or more green indicator lights or the red UNLKD light remain on. Action : Slow helicopter to 80 KIAS. Lower gear handle. Lf a safe down indication is obtained and destination can be reached with gear down, continue flight at airspeeds up to 130 KIAS. Do not consider raising gear again unless no safe landing area is within reach with gear extended. If gear will not extend, continue as in the section titled UNSAFE INDIC. TION - GEAR EXTENSION. ELECTRICAL SYSTEM FAILURES MASTER WARNING AND CAUTION-ADVISORY FAILURE With a latent failure of a diode in any of the caution or advisory lights that are on the test circuit, a subsequent failure in a corresponding system will illuminate all the lights on the warning or advisory s. By eliminating the electrical PO--wer to the failed system warning or advisory light, the lights can be returned to their normal condition. With the failed system identified, an appropriate reaction can then be made. Since a large percentage of the warning lights have been checked before flight by following the normal procedures section of the RFM, the following procedure is recommended for determining the most probable cause of the failure.

FAA APPROVED NOVEMBER 21,1978 Revised September 7,1990

Pan 1 Section 111 Emergency Procedures

K FLIGHT MANUAL

Proceed with each of the following steps, in order, until problem is eliminated. 1. Check engine, hydraulic, transmission, and fuel level instruments. If any abnormal indications are noted, pull associated warning light circuit breaker, and perform the appropriate emergency procedure. 2. Cycle OFF, if ON, these switches:

BLEED AIR PILOT PITOT HEAT CO-PILOT PITOT HEAT 3. Pull out and then reset the following circuit breakers:

DC ESSENTIAL BUS NO. 1 A1 ADVSY NO. 1 ENG CHIP &O. 2 ENG CHIP MGB CHIP I/TGB CHIP NO. 2 A1 ADVSY NO. 1 DC PRIMARY BUS BATT WAR& WSHLD HTR CONT

SO.2 DC PRIMARY BUS WSHLD HTR CONT NO. 1 FUEL LOW WARS (if indicated level is low). NO. 2 FUEL LOW WARS ((ifindicated level is low). DOOR WARK 4. Pull out both the following circuit breakers and then reset them:

DC ESSENTIAL BUS AC GEN WARN

NO. 2 DC PRIMARY BUS AC GEN WARN

FAA APPROVED KOVEMBER 2 1,1978 Revised September 7,1990

Pan I Section 111 SIK@ES~~~E Emergency Procedures WGHT MANUAL

BATTERY OVERTEMPERATURE Symptom:

BAIT HOT warning iight on.

Action:

Check B A n OFF caution list on. If caution i i a t is not on, place battery switch OFF.

I

Land as soon as practicable. If any smoke or odor is noted. land its soon as possible. SINGLE DC GENERATOR FAILURE

DC CiEN caution light on.

Symptom:

$1 or #2

Aaion :

Move affectedgenerator's switch to OFF. wait 5 seconds. move to RESET, then ON. Repeat as necessary.

If unable to'restore generator output. pla~xfailed generator switch OFF. If in IMC flight. transition to VMC as soon as prd~ricdble and continue flight. If a single generator fails. no loads are dropped The remaining generator will normaIly provide required power for all insdied equipment. Consider shutting off unnecessary equipment. however, as a conservative measure. The radio master switches. if installed. may be used to turn off non-essential avionics equipment.

FAA APPROVED NOVEMBER 21,1978 Revised March 2,1989

Part I Section 111 ~ I ~ B Emergency RMeduresLmGT M*Nu{

& H

~ /

'SA 4047-76-1

NOTE

The utility hoist will still function after a single generator failure, but the high speed will not be available. DUAL DC GENERATOR FAILURE

Symptoms: #1 and #2 DC GEN caution lights on. BUS .TIE caution light on - indicating that both dc primary buses are inoperative. #l A C PWR caution light on (Dual inverter system) indicating that No. 1 ac primary and 26 vac busses are'not powered (No, 1 inverter is inoperative).

NOTE At this time the only dc power source is the battery which powers the dc essential bas. Engine, transmission, hydraulic and fuel system instruments will be lost except for N,, N1, Tg , and fuel quantity. The pilot's fight instruments, navigation radios, and one VHF communication set will be retained. NOTE LOAD SHED SWITCH, if installed, should be given consideration. In the EMER LOAD SHED position, the engine overspeed system, transponder if installed, anti-collision light, and nowflight instrument dimming circuits are dropped from the DC essential bus.

I

Action:

Attempt to reset both generators by moving each generator switch to OFF, wait 5 seconds, move to RESET, then ON. Repeat as necessary.

If in IMC flight, transition to VMC as soon as possible. FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

SEEmZE FUGHT MANUAL

Part 1Procedures Section III Emergency

NOTE The utility hoist will be inoperative, but the hoist shear will still function with battery power. Reduce electrical load by shutting off unnecessary equipment. Land as soon as practicable. DC GENERATOR OVERTEMPERATURE - optional configuration Symptom:

#1 or #2 DC GEN HOT caution light illuminated.

Action :

Move affected DC generator switch to OFF. If DC GEN HOT light remains on, land as soon as practicable. If light remains on and indications of a fire are evident, such as burning odor or smoke, shut down the affected engine.

FAA APPROVED NOVEMBER 21,1978 Revised September 7, 1990

INVERTER FAILURE ( A C GENERATOR SYSTEM) (Applies to aircraft with INVERTER caution light installed) Symptom:

INVERTER caution light on.

Action :

If in IMC , transition to VMC as soon as practicable, and continue flight. NOTE Subsequent failure of the A C generator will cause loss of all A C electrical power.

AC GENERATOR FAILURE (AC GENERATOR SYSTEM)

Symptom:

AC GEN caution light on. light on.

Action :

Check that ac generator switch is ON.

3-18B

INV ON advisory

FAA APPROVED NOVEMBER 21, 1978 Revised July 10, 1987

F U M MANUAL

Turn ac generator switch OFF, then ON. If gene-r is not regained, ttvn switch OFF. ' If in IMC, transition to VMC as soon as practicable, and contirme night. NOTE The inverter is capable of providing ac power for all installed eq-ment except the optional heated windshield which i s on the monitored bus and will be chapped arrtomalicall~. .NOTE

On helicopters equipped witb AFCS Phase I l l , tbe AUTO SHUT DN light and loss of AFCS Pbase ID functions will be absemed as ac power source m?it&es £rum ac generator to inverter. Reengage by cycling cyclic trim to ON for 1 second and back to AUTO and cycling yaw trim to OFF for 1 second and back to ON.

-

AC GENERATOR Qxuptom: Action:

FAD.-

(AC GENERATOR m)

AC GEN BEARING caution light onC

~ Limitflightto5hoprs. ~

.

:The WINDSEiEUl IK)T caution light goes on when either the pilot's or 'copilot's windshield exceeds 58OC (135 OF). The system should be hnaed off to preveni over&& damage to the windshield If required, the individual overheated windshield can be identilied and left off to allow of the re5nainingwindshieldhalf. IfthecaationlightrePrainsonafterthewind-shield beater switches are turned of£, the weal cmtactors may have W d e d . The windshield power circuit bmakers should be pulled in this case. Sympbm:

WINDSHIELD HOT carrtion

Action:

Pilot and copilot windshield heater switches off WINDSHIELD HOT carrtion ligteL off D the caution light does not go off, pull .the PLT and LT WSHLD HTR PwR circuit breakers on the AC MON bns

Check:

Acti~n:

an

-

-

. If the heated windshield is reqPired and tbe power circuit breakers were not pulled, tke windshield can be restored as follows: Feel center of pilot's and copilot's windshield with hand, If one windshield feels dM&Uy cooler than the other, trPn an the windshield heater switch for the cooler windshield. If t h e W I N D S H i E I S H O T ~ l i g i ~ t r e t r r m txuntheswitch s, off and do not attempt to use either windshield heater. FAA APPROVED NOVEMBER 21, 1978 Revised MARW 24, 1981

EMERGENCY ENTRANCES AND EXITS

a

The h .doors are opened from the oatside by Lifting the door fatch or framinsidebyl~qthel~~mthelow~affarindow~eof each door and pplling the door latches aft. The doors m a y also be jettisoned from inside using the door lockpin and jettison bridles, forward of the door latches. The jettisaDl levers are under a plastic protective . A fmger grip in the plastic cwer is used to pop the cover loose from the door fame. The red jettison handle is pulled up to release the hinges, and a sharp blow with the £idan the door just aft of the door handle will jettison the door. Decals as shown in Figure 3-2 provide for normal opening or jettiganing af doors.

IIf door does not open normally, attempt to jettison it. Door Jettison Procedure

1.

liftdoorlockpin

2. Remove jettisan handle plastic cover 3. 4.

P\rlljeUisonkandleup door jusL aft of the jettisan handle sktrply with the dist

RH. S I m r n G DOOR (OP?IONAL) The R H 8liding door is opened from the astside by docking the door lo& (raS;I.tingthe door locktothe UNLOCgwm) and then mbihgthe dour handle downward to tke OPEN positictn. Pull the door orst and slide aft. Fhxnthe-e, mwethedoorlocktotheUNLOCgpositionandthen ruktethedoorbandledownwardtotheOPENpositi~ Pushthedoora& a& slide aft. If the doar does not open normally attempt b jettison - 0 1 ~ . Windaw Jettison Procedure

NOT33

Do not jettison window in flight 1. ~e~plastictoverovertkwindowjettisanhrmbir. 2.

Pullwindow jettisonhandle.

3.

Rlshwinduwarrt.

FAA APPROVED NOVEMBER 21,1978 Revised July 26, 1982

PlLQTSDOOR

*

ENGER W O R

*

m l D E ENTRAUCE UIT DOOR HANDLE AT Am END OF EACH DOOR

INSIDEDUT

1

Figare 3-2. Emergency -es

and Exjsts (Sheet 1 of 2) FAA APPROVED NOVEMBER 21, 1978 Revised APRIL 12, 1979

3-20A

DO NOT OPEN

PUSn HERE TO CHECK THAT DOOR IS LATCHED

APPUCABLE TO RIGHT SDE DOORS NOT MODIFIED BY PIN 7607O200L5911

AND AU LEFT DOORS

-3-2.

Bne%en~cesandEXts(Sheet2~f2) FAA APPROVED NO-ER 21, 1978 Revised OCTOBER 14, 1980

Right Sliding Door Not Modiiied by Kit P/N 76070-20015-011 aad al3 Left Sliding Doors Do not o-p& the door in fight.

In-Flight Door Uplocked Procedure DOOR OPEN caution light on. Symptom: Action:

1. Reduce to and maintain minimum safe speed. 2.

Turn seat belt switch on.

3. If crewman is available, have crewman check cabin sliding door to insure closed. If closed, DOOR OPEN caution light system has malfpnctioned. If door is ajar, land as soon as practicable. Have crewman hold/puU door handle inward to prevent from opening outward and sliding aft. In-Flight Door Open Procedure Symptom:

DOOR OPEN caution light on and door visibly ajar.

Action:

1. Reduce to and maintain minimum safe speed. 2.

Turn seat belt switch an.

3.

Land as soon as practicable. Avoid abrupt maneuvers.

4.

If crewman available, bave crewman hold/pP11 door irrraard to prevent from slidirrg aft.

5.

Ifdoorcannotbeheldinplace, m a s soanas possible.

Rigfrt Sliding Door Modified by

git

P/N 76070-20015-011

If slichng door becomes unlocked or opens in flighf decrease airspeed to 50 KXAS o r less before closing. If not possible to close door, open fully and lock. Do not exceed 74 gIAS for remainder of flight.

FAA APPROVED NOVEMBER 21, 1978 Revised OCTOBER 14, 1980

I

The optional emergency flotation system is designed anly for emergency landing on water. The system may be eqected to keep the helicopter upright long enough to permit engers and crew to exit to life rafts or rescue boats. A subsequent takeoff o r lung term towing should not be attempted.

I

When it anpears that d i t c u is likely, coasideration must be given to such factors as wind direction and velocity, sea state conditions, and aircraft power available for ditching. The maxim~m) permissible water e A r y conditions are 33 kwts water speed a t 300 feet per minute rate of descent in a calm sea. Optimum ditching conditions would occur in a calm sea state with the forward speed of the helicopter reduced to as near zero as possible, and with little or no lateral drift component. Minimum touchdown forces will be achieved whm touchdown is made on the crest o r back of a wave with a minimum ate of descent. Greatly increased touchdom forces will be experienced i f the fanding is made on the front o r rising face of a wave. Every effort should be made to land the helico&r with as little sideward drift as possible as the roll rate after touchdown increases sharply with any increase in lateral motion. Ditching with power available (suchas when fuel starvation is imminent, or aloss of transmission oil pressure dictates sach action) will greatly increase cantrollabiliQ, touchdown forces, and assist in p*venting the helicopter fram rolling after impact. Power off ditching (autorotation) should be avoided i f possible. With sufficient power available to frilly control the aircaft descent rate, sideward drift and forward speed to near zero values, successful ditcbbgs may be accomplished in sea states v q to aad including Sea State 4 (wave height 6.5 feet, wave leagth to height ratio - 1to 10) depending on wind conditions. All possible contrdl -e from the rotor system shonld be used to prevent rolling after impact. Consideration shaPld be given to extended power on wakr taxi, if wave conditions make rotor to water improbable. Power-on water taxi will greatfy increase the roll stability of the aircaft and will allow yaw and heading c-. The f l h may be hfhted at any time daring the ditching procet3ure (below 75 KEAS). Wow SrrfEiaent time for full hfklion before water confad. Float innatiaa-time is within 10 seconds. The following genead dikhbg are recommeaded as the best for a successful controlled emergency m r landing:

~ ~ & W o n s p s t e m i n a h o e r i f ~ c i m ~ r i s a ~ to permit a hover. Reduce forward speed to as near zero as possible. m e lateral drift companent to a minimum. Avoid rising faces of large waves, ReduceraieofdeseenttominimmnattcroE~. Fuselage angle at impact should be between O 0 and lo0 nose ap. UseaUcantrolavaifablefromtherobrsystemtop=veatthekelicopterfrompitchiPgorroIlingafter~wn. The pilot should endeavor t o touchdown, heading into wind, as near the crest of a wave as possible to minimiffe the impact vertical velocity. Ifthehelicopteristowed, itshoaidbetuwedataslowmhof speed. FAA APPROVED NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

1. natation -stem test switch 2. Landing gear

- Check NORM.

- DOWN. NOTE

If the landing gear cannot be e d e d for any reason, the floats can be dephyed with the gear reh-acted. CAUTION

Operation of the landing gear during or after float inftatian may result in damage to the floats and parkid loss af flotation capability.

4. Press the cyclic stick float innntion switch. Allow sufficient time for

full innatinn prior to water .

NOTE

I

noat innltinn time is 10 seconds.

-

5. Notify all occupants to prepare for emergency water landing.

NOTE

Cabin occslpants should h s k n their safety belts and remain in their seats with their anns against their knees~watercantacti~made.

NOTE

H time and conditions permit, consider opening and locking r. h sliding doar if m a e d by Eit P/N 76076-20015-01L 6. Tansmit distress message,

7. Check life vests for seritp and lock sbodlder harness. 8. aald the heliwpter level as possible, con&& forward speed and rate of descent.

the water with minimmn

9. Water and aircaft conditions permittirrg-consider extended water taxi.

FAA APPBDVED NOVEMBER 21, 1978 Revised JANUARY 17, 1983

10. If the helicopter is unstable or tatdng on water, shut &m engines, notify cabin occrrpants to evactlate through the cabii doors after the rotor blades have stopped turning. Use of the rotor brake is not recommended, especially above 40%Nr.

Do not evacuate the helicopter anti1 the rotor blades have stopped turning. Do not Slate life vests until clear of the helicopter.

FAA APPROVED NOVEMBER2i, 1978 Revised July 26, 1982

I

Part 1 Section m Emergency Procedures

R l M MANUAL

EXTERNAL CARGO HOOK PROCEDURES

Tf an emergency requires release of the external cargo, perform either of the following:

1.

CARGOFiOOK~&t&-ARhdED

2.

CARGO hook release button (pilot's or co-pilots cyclic stick)DEPRESS.

1.

EMERGENCY CARGO HOOK RELEASE 'T' bandle (on center console)-PULL

Because of the rapid CG shift following a sudden release, a moderate climb and pitch rrp should be anticipated.

NOTE

If an approach terminates with bd£icient power to stop the rate of descent, consideration s M d be given to allowing the cargo to touch down ratber than executing a premature release. UTIUTY HOIST PROCEDURES

1 Xf an emergency requires release of a utility

hoist load, perform the

following:

1.

-

Cable shear switch @ifutts or cabin) SHEAR. I£ load does nut separate, CIEtthecablewithtke~cablecuttersasclosetothehoist as possible.

For hoist opewith electrical ma3£nnctions, refer to ELECTRICAL SYSTEM FAILURES in this Section.

FAA APPROVED NOVEMBER 21, 1978 M e d JUNE 6,1980

Part 1, Section IV Performance Information

FLIGHT MANUAL

TEMPORARY REVISION NO.2k

4

FILING INSTRUCTIONS: Insert tMs Temporary Revision facing page 4-1, dated July 26,1982. SUBJECT: CATEGORY "A" AND 'B" OPERATION4

This e tr a M i adds a rejected takeoff and landing d i m adjustment for aitcraft !hat have 5014067 rotating brake disk($) instded as determined pm Aircraft Braking Systems'Alert Service Bulktin No. S76-32-A24 or S76-32425. MANUAL CHANGFS: If the 5014067 h k e eliks are linsbUedor the brake disk configuration has not been confirmed, calculate the Category "A" Rejeded Takeoff Distance end Category "A" Larding Distance or Category "B" Landing Distance normally, as appHcafAe. Multiply the distance@)(by 1.67 to obtain the correcbd distance(s).

,

Figure Number 4-8

Ti Category -AwRejmed Takeoff Distance

1

4-14

I

4-19

1 Category 'B" Landing'D i n c e

Category "A"Landing Distance

FAA APPROVED APRIL 9,2002

SECTION m

Density Altitade Chart..

..............

- Pilot's System.. n - Copilot's Spstem .

4-3

Aizspeed cafibration

4-5

Airspeed m

4-6

Torque Available, 2-1/2 Minute Power OEI and 30 Mimte Power OEI.. ......

4-8

Totqoe A e l e , 2-1/2 Minrrte Power

OE3 and 30 Mimte OEI, both with EAPS Installed and Switched OFT

....

Torque Available, Takeoff Power Dual Engine, and Normal Crmise Power DualEngine

........................

4-8A

40B

Torque Available, Tag& Power Dual Engine, and Normal Cmi8e Power, both with EAPS Ins&lled and Switched

ON ................................

4-9

........

4-10

Category "A" TakeoE Profile-.

~ " ~ " ~ D i 8 t a n C e ~ AehigveTageoffSafety Speed 4-22

-

........

Foraard nimh Performance Single Engine, 2-112 Minute Pawer, !bk&Z -Speed.

.......................

4-14

FomardrrrimhPerformabce-5gle Engine, 3OMimte m r , Bestof CIiimb Speed

..................... 4-15

ForrP;udrrrimhPezformance-SEngle Engine, Maximum Conmnms Power, Bestof Climb Speed..,..,,.,,,,

FAA APPBDVED

416

NQVEMBER 21, I978

Revised July 26, 1982

4-1

mm

PAGE

Categorg "A" Landing Distance

Category

""

T k d e a f f Distance

........

4-18

........

4-19

.... Forward Qimb Performance - Two

4-20

...

4-21

..........

4 -22

F o W Climb Perfoxmace Two l h g g s , Takeoff Power, 52

Errgins,, m r r m C~~ Power, Best Rate of Climb Speed..

-

Forward Qimb Performance Z b o Engines, Normal Cruise Power, Best Rate of Climb Speed..

...... Altimeter Calibation - Pilot's and Copilot's Spstem,. .................. Category "B" Lauding Distaace..

Bover Out of Ground Effect

...........

FAA APPROVED NOOEMBER 21, 1978 Rwisea July 26, I982

4-23 4-24 4-25

Part 1 Section IV Performance Ipformatim

NGHT MANUAL

FAA APPROVED NOVEMBER 21, 1978 %vised JANUARY 9, 19'79

FLIGHT MANUAL

FAA APPROVED NOVEMBER 21, 1978 Revised JANUARY 9, 1979

AIRSPEED CAUBRATION PILOT SYSTEM POSITION ERROR ONLY AERO MECHANISM PART NO. 8 u m ; s 2 0 L W OR AEROSONIC PART NO. 20020-11190

AIRSPEED lNDICATOR

Figure 4 3 FAA AP-VED

NO-

Revised 3uly 26, 1982

21, 1978

AIRSPEED CALlBRATlON COPILOT SYSTEM P o s m o N ERROR ONLY AERO MECHINlSYlPART NO. BWltCS20LW OR A E R o s m I C PAF5 NO. 2002&11190 AIRSPEED lNDtCATOR

Figtae 4-4

FAA APPROVED NOVEMHER 2 1 1978 Revised Jrily 26, 1982

TORQUE AVAILABLE

Torque available charts are provided in Figures 4 5 through 4 6 A to determine the engine torque available at various pressure altitudes and oatside air tempemtnres. Each chart is based on the specific engine power rating and forward speed. As noted on the charts, Figures 4 5 and 4 6 a r e for helicopters with EAPS not installed. Figures 4 5 A and 4 6 A are for helicorkers with U P S irrs&lled with switch wsition (ON or OFF) as specified on each cSuperimposed on these &s a r e t& engine m&dac&er's high temperature engine operating limit and the engine o r aircraft transmis sion torqae limit as applicable. The engine and aircraft transmission torque limits a r e listed as follows:

ENGINE

TRANsIassxON

TORQUE

TORQUE LIMITS-%

POWER RATINGS

L ~ M ~ T S? - i

f

f"1D Minute pwer

*

THIS VALUE RELATES TO RECO-TIONS MANUFACTURER SEE PART 2, SECTION L

OF THE EN-

In the instances where engine torque limit otceeds the transmission t o m limit, then the lower a e is used in determining aircraft performance. Wer to Pigure 1-5, "Engine or DrWe Spstem Opemthg Limits for compliance with m.

Unless othe-e noted, the performanee data shown on the charts is for helicwithout EAPS installed. If the EAPS -on (with switch either ON or OFF) affects performance, the decrease in performiince is noted an each char&

FAA APPROVED NOVEMBER 21, 1978 Revised NOVEMBER 2, 1981

46A

ROTORCRAFTEXTFZNAL LOAD CAWGORY "B" OPERATIONS When external load operations are being conducted, the following performance factors should be .cansidered: 1.

The helicopter was evaluated under day and night VTR conditions for diiiQ hoist loads and day VFR conditions for cargo hook load. This does not preclnde the operator from demanstrating the a b w b operate rmder night VFR conditions for cargo hook loads.

2.

Performance published m the w t manual may or may not be realized, depending on the size and shape of the laad Coqensatian

should be allowed for the effects of external loads having a high dag, particularly in those performance areas involving forward speed, such as takeoff distance. 3.

Flight checks were performed to establish safe operatian for Class *'B" external 1osLds as'follows:

a.

kn The cargo hook was tested nsiag high density loads weight from 2600 pounds to 3300 pounds. The tests were conducted at densitj altitudes of 1000 feet, 3000 feef and 9800 feet.

b.

The utility hoist system was tested with a 600 p o d load coosisting of an apptorcimatefy 20 inch cube of cement at various cable lengths to a Vne of 74 E;IAS. A low density of 238 pounds Load consisting of an open top box, measm3ng 40X40X37 inches was tested to a Vne of 57 KIAS. A 155 pound simulated person was hoisted in a bover to the aircraft and bmughi into the cabin, Tests were conducted at 1700 feet density altitude and below.

Grarmd kandling -cs were evabbd during representative maneuvers at maximum gross weight an a bard surface rrmway.

FAA APPROVED NOVEMBER 21, 1978

Re-

NOVEMBER 2, 1981

I

FLIGHT MANUAL

TORQUE AVAILABLE BLEEDIlR OFF ENGmE ANTEICE OFF

-

EaPsmrr1NsrAuEu GENERATOR LOAD

200 AMPS

RATING: 2-In MINUTE OEl 96% Nr 52ENGINE TORQUE UMI=1112 %

RATING: 30 WNUTE O n 96%Nr BROCSPEEO LUIN TRAWSYlSSlON TORQUE UYIT=lO(%

Figare 4-5 FAA APPROVED NOVEMBER 21, 1978 RePised UNITARY I?, 1983

Part 1.Seetian N

S4 4047-76-1 FLIGHT MANUAL

TORQUE AVAILABLE ENGIUE ANTLICE OFF

BEEBAlR O R

GENERATQR L O A M AMPS RATING 2-V2 MINW'E OEl 5210AS 96%Nr EN6lNE TOROUE UldTT=llU%

40-30-20-lo

0

1

0

2

0

3

0

4

0

~

OLrmDE AIR TEMPERATURE-.C

I

Figure 4 5 A

FAA APPROVED NOVEMBER 21, 1978 Revised JANUARY 17, 1983

48A

TORQUE AVASLABLE B

W

R OR ENGINE AKFI.lff UPS NOT INSTAUED

GENERATOR LOAD-

OFF

200AMPS

RATINP TAKEOFF DUAL-ENGINE IGE 0 KUS 10@% Nr

Figure 4-6

FAA APPROtTED NOVEMBER 21, 1978 Re-d JANUARY 17, 1983

TORQUE AVAILABLE BLEEDdlR O f f

ENGINE ANTI-ICE O R

GENERATOR WADr200 AMPS

RAf)NG TAKEOFF DUAL-ENGINE IGE 0 KlAS lW% N r

EAPS INSIALlED AND SWITCH ON MAIN 7RANSMISSION TOROUE UMCT = 100%

io-jo-50-ib

o

i o z b i o * o s i

OUTSIDE AIR TEMPERATURE-'C RATlNG NORMAL CRUSE DUAL EN6lNE BROC SPEED 100% N I EAPS~AUEDANDSmTCH OFF ENCIWETDRQUE R A M = 863 %

OUIStDE AIR TEMPPUTURf-.C

Figure 4-6A FAA APPROVED NOVEMBER 21, 1978 RePised JANUARY 17, 1983

fl 4-9

FAA APPROVED NOVEMBER 21, 1978 Revised AUGUST 27, 1979

-

FLIGHT MANUAL

Figure 4-8 FAA APPROVED NOVEMBER 21, 1978 Revised MARCH 23, 1981

FLIGHT MANUAL

FAA APPROVED NOVEMBER 21, 1978 Revised M C H 23,.1981

FUGHT MANUAL

FORWARD CLIMB PERFORMANCE The climb charts (Figures 4-10 Wet@ 4-12 and 4-16 thraugh 4-18) provide

a means of detemPining rate of climb for combinations of gross weight, OAT, and pressure altitpde. Each chart is based on a specific power rating and forward speed To determjae helicopter rate of climb, enter the appropriate chari at ihe ambient aatside air temperatpre, and move vertid y upward to intersect the iiight pressure altitude. From this point move horizonb.Uy to the right t o intersect the flight gross weight, then vertically downward to read rate of climb for the given fiight conditios.

FAA APPROVED NOVEEMBER 21, 1978 Revised NOVEMBER 2, 1981

I

Figure 4-10 FAA APPROVED NOVEMBER 21, 1978 . Revised MAY 14, 1985

FLIGHT MANUAL

rigme 411

FAA APPROVED NOVElVlBER 21, 1978 Revised MAY 14, 1985

FORWARD CLIMB PERFORMANCE SINOLE.EN0INE MAUIMUM CONTINUOUS POWER BESTRATEOF CLIMBSPEED 96% Nr ANT 1-ICEOff NO BLEEOaAIR OEAR RETRACTED

REDUCE RATE OF CLIMB D n t R M l N t D FROM CHART BY AMOUNT SHOWN I N TABLEI ROC OROS8 REDUCTION WEIOH EAPS -109 INSTALLED '

5700 6000 7000 8000 9000 10000 10500

45 45 35 30 30 25

ROC REOUCTION ANWCE ON

90 90 80 75

70

ROC REDUCTION PAPS INST ANTl4CE ON

175 170 180 140 140

ROC REDUCTION UTlllTY HOIST INS1

40 40 40 35 30

ROC REDUCTION RIQHT SLIDINQ DOOR OPEN

AIR TEMPERATURE-@C

t

15

115 110 100 95 90

-

OROSS WLtOlIt 109 (KO) X 1000

9 ARE CUMULATtVE

-

RATE OF CLIMB 100 FPM

BROC SPEED

O' 18

a

l2

I 13

4

W

25

W I M COMPOUND CONfIOURAT

t

=

O

FAA APPROVED NOVEMI(ER 21, 1978 Revised JANUARY 17, 1983

Figore 4 1 4 FAA APPROVED NOVEMBER 21, 1978 Revised JANUARY I?, 1983

CATEOORY "8" TAMEOFF DISTANCE FROM 8 . n HOVER TO 80 FT

D18TANCLI ARE BASED ON W E S t CONDITIONS: 1. MAIIlMUM ALLOWABLE WEIOIW 2. WIND.CALM 3. NO BLEED.AIR USE EKCEPT ANTI-ICE 4. ANTI4CE OfF OR O N

EIIAMPLE: PRESSURE AL11TUDE- 1,000 M E 1 AIR IEMPERAWRE-2I'C f A i ( ~ 0 1 l DISTANCE- 1,040 fEET

FORWARD CLIMB PERFORMANCE I W O ENQINES TAWEOIF POWER 52 MIA9 100% Nr N O B L E E D W ANII4CE 011 OEAR REIRACIED

REOUCE RATE O t C O M B DtlERMINED FAOM C t I A R I BY A M O U N I SHOWN IN IABLE: ROC ROC ROC anoss RCOUCTION REDUCIION REDUCTION WEIO~II EAPS ANT~.ICE EAPS INSI LBS INSIAUED ON ANII~ICE ON

-

6700 6000 7000 8000 9000 10000 LO500

105 105 100 I00 95 90 90

.

120 115 I10 105 105 100

220 210 210 209 205 190

ROC REOUCIION UTIL~TY IIOIS~ INSI

40

ROC REOUCIION RIO~II SLIDINO DOOR OPEN

115

40

110

40 35 30 30

100 95

90 05 100 190 30 85 W l t l l COMPOUND CONIIOURATIONS. RATE Or CLIMB CORRECIIONS ARE CUMULATIVE

AIR TEMPERA7URE-4

*

-

OR059 W E l O l l l LBS (nab n 1000

an

Y

RATE OF CLIMB* 100 FPM

S S6WJ (C33)

FAA APPROVED NOVEMBER 21, 1918 Revised MAY 14, 1985

F U M MANUAL

Figure 4-18 FAA APPROVED-N Revised MAY 14, 1985

21, 1978

CATEGORY "B" U N D I N I DISTANCE FROM 80-FOOT HEIOHT TO STOP ONE ENO~UE luormttvt

ALTIMETER CALIBRATION SEA c

m c STD tamp

NOTE: DOES NOT INCLUDE INSTRUMENT ERROR

5 %

"O

9"O

PILOTS SYSTEM

-. .

0

20

40

60

80

100

INOfCllttO MI) 9 0 E L b 4 N O T 3

HOVER OUT OF GROUND EFFECT TAKEOFf PO100% N

,

ANT&-ICEOFF

NO -AIR

REWQ 6ROSS WEK3fT 0-W I ED FROM CHART BY AMOUW SHOWN IN THE FOLLOWING TABLE. AS APWCABLE: WEIGHT REDUCTION 150 POUNDS 200 POUNDS 270 POUNDS

CONFlGURATlON E4PS lNSTAlLEO a ANTCICE ON ANTClCE ON WITH UPS l N s A U E b

.

io

.

CROSS WEIGHT-lo00 KUOCRIlYtS

i'

-

sb

s - 0

rn

4-21 FAA APPROVED NOVEMBER 21, 1978 Revised MAY 14, 1985

4-25

SA4047-76-1 Supplement No. 32

*

FLIGHT MANUAL

ERRATA

To correct a printing error, replace the Title and Introduction page, and add page iAi to the manual supplement.

Supplement Index

FUGM MANUAL S-76A SUPPLEMENT INDEX

The supplements listed below are not applicable to every serial number helicopter. The operator is only required to have those supplements that pertain to the equipment installed on his helicopter. If the equipment described in a supplement is installed by the operator at a later date, the applicable supplement may be ordered from Sikorsky Aircraft. Annotate with an X in APPLICABLE column those supplements that apply to your helicopter. Please update any revised Supplement Index, as required, when received with future changes. SUPP NO.

APPLICABLE

1

-

2

-

SUBJECT

DATE

Collins ANS-31 RNAV, MLS receivers, DME, and Com Systems for helicopter S/N 76OO7.

1014f79

ADI, HSI, and BEARING selectors,

10/26/79

FLT director and HIGH FREQUENCY RADIO for helicopters SM 760021, 760022, 760024, 760025, 760026, 760039, 760046, 760052, 760056, and 760057. ADI, HSI, and Bearing Selectors, FLT director and HIGH FREQUENCY RADIO for helicopters S/N 760036 and 760049.

2/8/80

GNS-5OOA OmegaNLF for helicopters SIN 760012, 760013, and 760119.

5113181

ANS-3 1A RNAV, DME40, an C-14A Com Systems for helicopter sm 760065.

6/6/80

Navigation Switchin2 Equipment for helicopter SIN 76OO62.

7/3/80

NOVEMBER 21, 1978 Revised June 17, 1997

I

Supplement Index FLIGHT MANUAL

S-76A SUPPLEMENT lNJ3EX SUPP NO.

APPLICABLE

SUBJECT ANS-3 1A RNAV, Dual D M E 4 and C-14A Com System for helicopters SIN 76007 1 and 760074. Navigation Switching Equipment and TDL-711 LORAN C for helicopter SIN 760081. Litter Installation Navigation Switching Equipment and TDL-7 11 LOWS!? C for helicopter SIN 760116 Navigation Switching Equipment for helicopters SIN 760 101, 760 102, 760130, and 76131. GNS-5OOA OrnegaNLF Navigation Switching System for helicopter SIN 760133. GNS-5OOA (Series 3) Navigation Switching System for helicopters SIN 760159, 760161,760176, and 760185. Sperry FZ-702 Flight Director for helicopters Sm 760159, 760161,760176, 760185, 760189, and 760190. AYS-3 1A RNAV, Dual DME,4, and Dual C-14A Com Systems. Dual HSI Control Switching with Flight Director ANS-31A RNAV, Dual DME-40 and Single C-14A Com System Navigation Switching Equipment, TDL-7 I I L O W - C. and UHF DF.

NOVEMBER 21, 1978 Revised June 17, 1997

DATE

Supplement Index -

FUGHT MANUAL

S-76A SUPPLEMENT INDEX SUPP NO.

APPLICABLE

SUBJECT

DATE

19

Global Navigation GNS-5OOA (Series 3) Installation

1/12/82

20

Sperry FZ-702 Flight Director

1/12/82

21

SIM - Spindle Inspection Method.

12/27/83

22

Navigation Switching Equipment and TDL-7 1 1A LORAN C for helicopters SIN 760238, 760243, 760266, 760267, 760277, 760278, 760279, and 760280.

513 1/84

23

Auxiliary Essential Power System.

312 1/85

24

Navigation Equipment Switching System, TDL-7 11A LORAN C, and LTN3000 OMEGAIVLFIRNAV for helicopters SIN 760275 and 760276.

7/6/84

GNS-500A (Series 3B) Navigation Switching System with pushbutton selector switch .

12118/84

Low speed approach procedures for microwave landing system and ILS.

3/27/86

Placard for Aerosonics PIN 2002011190 airspeed indicators.

1/31/86

Tail Rotor Servo Shutoff Valve

12/12/86

Turbomeca Aniel 1 s engine installation

6117/97

Turbomeca Aniel 1S 1 engine installation

12/17/97

1

Turbomeca Aniel 1S 1 engine installation (Increased performance credit)

12/17/97

1

DC Essential Bus Recovery System

8/25/88

NOVEMBER 21, 1978 Revised May 15, 1998

Supplement Index

F U G M MANUAL

S-76A SUPPLEMENT INDEX SUPP NO.

-

APPLICABLE

31

Alternate Static Source

33

-

34

-

Bendix RDR- l4OOC Weather Radar

35

-

Not printed

36

-

Snow Landing System

37

-

AEM Loud Hailer

32

38

SUBJECT

Dual Static Inverter Metair Life Raft Deployment System

Category " B Operations With 9 or Less enger Seats

SOUTHERN FAA REGION SUPPLEMENTS S- 1

-

RCA Data Nav 111 System

S-2

-

GNS-5OOA OmegdVLF Navigation System

S-3

Left-Hand Sliding Cabin Door

S-4

-

S-5

-

Cyclic Stick Guard

S-6

-

Collins ANS-3 1C RNAV System

S-7

-

Speny ZC-301 Flight Director and 7000466-936 Attitude Indicator

S-8

-

Canadian Marconi CMA77 1 Omega1 VLF Navigation System

S-9

-

Collins ANS-3 1C RNAV System

S-10

-

Long Range Fuel System (106 Gal)

4

Bendix ILJ2023 and CC-2024 Navigation Equipment

NOVEMBER 21, 1978 Revised May 15, 1998

DATE

Supplement Index FLIGHT MANUAL

S-76A SUPPLEMENT INDEX SUPP NO.

APPLICABLE

SUBJECT Foster Airdata Systems LNS-6 16A Long Range Navigation System UNS- 1 Jr. Navigation System KNS-660 Flight Management System Bendix MU-21 Microwave Landing System Not printed Metair Long Range Fuel System (PIN 33776-6 1049-013 or 33776-61049-014) (50 Gal) UNS- 1 Jr. Navigation System Bendix Aerospace RDS-86 Weather Radar System Bendix Aerospace RDS-84 Weather Radar System UNS- 1A Navigation System Not printed Honeywell EDZ-705 Electronic Flight Instrument System P-870 Weather Radar System Honeywell LSZ-850 Lighting Sensor System Dual Honeywell SRZ-850 Inteagated Radio System Cabin Lighting Control System

NOVEMBER 21, 1978 Revised May 15, 1998

DATE

Supplement Index

S-76A SUPPLEMENT INDEX SUPP NO.

SUBJECT

DATE

Cabin Environmental Control System (ECS)

10/7/94

Global Wulfsberg System Flitefone VI

11/28/88

Grimes Recognition Lights

11/28/88

ARNAV ELT- 100HM

12121/89

Pulselite System

11/28/88

GNS-500 Navigation System

3/22/89

Not printed Emergency Medical System (EMS) Interior

4120189

UNS- 1A Compact Flight Management System

51919 1

Spectrolab Nightsun Searchlight System

4/27/90

Forward Looking Infrared (FLIR) System

4/27/90

SPZ-7600 Digital Automatic Flight Control System

6/4/92

Not printed Collins DF-30 1E

10/11/90

Not printed Collins Dual or Single ALT-55B Radio Altimeter System

4119/90

Baker CockpitICabin Audio System

4/27/90

Not printed NOVEMBER 2 1, 1978 Revised May 15, 1998

Supplement Index

FLIGHT MANUAL S-76A SUPPLEMENT INDEX SUPP NO.

SUBJECT

APPLICABLE

Honeywell Primus P-7001701 Radar

DATE 6116/92

Not printed RACAL RNAV2 Navigation Management System

10/18/90

Not printed Not printed Not printed Radio Altitude Warning System (RAWS)

4/23/90

Honeywell Data Nav 111

10111/90

Logo Light System

9/8/89

External Sliding Door

6/23/94

Cargo Hook Load Indicating System

4/27/90

Variable Speed Utility Hoist

2/7/97

Battery Feeder Fault

5/30/90

Bendix EFS-10 Five Tube EFIS System

7/30/90

Not printed UNS-1B Multi-Mission Management System

12/1/97

1

Not printed LITEF LCR-92s AHRS

4130197

BendidKing CAS-66A TCAS 1 Traffic Collision Avoidance System

3/30/98

NOVEMBER 21, 1978 Revised May 15, 1998

1

Supplement Index F U G M MANUAL

S-76A SUPPLEMENT INDEX SUPP NO.

APPLICABLE

SUBJECT

S-64

-

P&G Solid State Combined Voice and Flight Data Recorder

5/7/96

S-65

-

Trimble 210 1 LIO Approach Navigator System

220197

NOVEMBER 21, 1978 Revised May 15, 1998

DATE

FAA APPROVED

ROTORCRAFT

FLIGHT MANUAL SUPPLEMENT

.REVISION NO. 1

SIKORSKY

MODEL

PART 7 LONG RANGE FUEL SYSTEM

Date aP Apptoval: f EBRUARY 24,1984

Date of Revision: January 30, 1995 Federal Aviation istfation

Atlanta, 'Georgia

rm7&' MANUAL FLIGHT

PART 1 LIST OF IFFEtC'TlW. PAGES FAA ROTORCRAFT FLIGHT MANUAL

FAA A P P R O W FEBRUARY 24,1984 Revised January 30,1995

Part 1 List of IilYtutivc Pag-Skippltsm~ntS-10

PART 1 LOG OF REVISIONS

The revision d a t ~is shown at the boftom of each revisad p a g oA vrtr~bnlbar cm the outer margin indicatff the latess revised porcion of each page. A vertioal, bar next to the page number indic4tes the page has bem retyped with no significant dmges, REVISION

PAGES

NO.

REVISED

1

REMARKS

DATE

FAA APPROVED

Added optional 1/39/95 aw tank (forward) hstallatim Changed title ,mi refomfitted

1-1, 1-2,2-1. 3-2,3-113-2

,

supplement.

m Manager, Aia&

Certification Oflice, FAA, Atlanta, 5 A

FA4 APPROVED FEBRUARY 24,1984 keviwed Jaauary 30,1995

i

Par1 1

Introduction

Supklon~cnl3- 1b

FLIGHT MANUAL INTRODUCTION

This dupplen~entmust be attached to the Sikorsky S-76A, FAA approved, Ratormatt Flight Mmual, dated Nov~rnber2 1,1978, or later FAA approved revision, when a Hunting Aviatiou/Mctair Long Range Fuel Sysrern (P/N 14-54) is installed in iucordanec: with Drawiq No,- 33776-61013 per STC SW157dSO. The informatio~loontained herein supplements the information of the Rotoror& Flight Manual. For limitations, procedures, and pd~&ce h & n not contained in rhis supplement, consult

the Rotorwaft Flight Maual,

FM APPROVE0 FEBRUARY 24,1984 Revised J m u 30,1995 ~

OPERATING LIMITATIONS

L

The aux lel,tauk shall be in&tdled and removed in insacordance with the Hunting AviatiodMGt(LiTirhtanmceManuid, Apply forward or aft instaUstion plaoard m r e q ~ Fud ~omectionsshould be examined for I d c a p after instaliationof ta&

I

WEIGHT AND CG W T S

" I ~helicopter Q whm fittad with the mdiary h 1system, is to be operafedwithin tbe weight aad center of gravity limits ypecijiod in Part 1 ofthe S-76A Rotor~aftFlight Manual. PartL;ular attention should be paid tb the embr of gravity oalculation before each flight, Rdw to the placard on thc insida of the baggage; c o m p ~d t~ o rto establish whether the LK& is oquippkd with a forward or aft auxiliq tank installation and then use the proper charts in the LOADING INOXMATION, Part 2,Section 1#ofthis suppletnent

i

W h ~ nthe aw fuel system is ki!italIed and the awx mik is m use, no baggage, hazardous, or flmmabIe items m y be cmied in the baggage cornpastmat. The baggage compubnent should be placarded W O 73AGGAO.E TO BE C m W N AUX FUEL TANK IS FITTED."

The aux tatllr is u, be gravity fueled only with fie1 whose water content has bm obecked by a

suirable metbod.

Fueling is ody to be c~tri,edout with the helicopter in a normal ground (lwel) a a i W t~ ensure adequate a p a m i ~ nspace in the nux tank, The

fuel vdve must be shut d&g takefl and lmdkrg to enswe fuel system m~endmce,

la-flkhtW tmsfkrfr~rnthe a w tank, to the main tadksmay be accomplidud in ~ d sflight, e stmiat and level, h m g transfer of fuel from the aux rmk UIthe main tanks, the aux valve~shouldba left io the O?EN position for at lwf 15 minutes after the awc h l g ~ g ehas switched fim I20 pounds to zero. This is to make sure h t all usable fie1 has transferrred.

FAA APPROVED FEBRUARY 24,1984 Revised January 30,1995

-.-

Pa11 1, Section I

Operating Tinitations Supp!ement S- 10

Whenfur;.lis tt.&etred

FLIGHT MANUAL

in siaight wd level aruise flight, them i&ma u n W l e fuel in the aux

taok. p T ~\TS~JU,T,ATZON: (See p h d iuside baggage ootnymmt door tcr identify iastrrlkation,) Four g d m s ifs thma-um alnouat of d l 8 fuel that raains in the awdiary tank, irad This rn-urn mo110.t 'kcmiwhen fudis transferred d 155 knots. Fuel transfer at the higher nose attitudes associated witb cruise speeds lower k m 155 knots reduecs the unusable fuel to less than the 4-gallon maximurn.

FAA APPROVED FEBRUARY 24,1984 Revbed .hnuary 30,1995

SECTION fl NORMAL PROCEDURES

EXTENOR INSPECTION When the a.ux tMk is fitted and in use, before thc first flight ofthe day, or bdbo1ethe next flight aAer extendcd maintmwm, perform the following p r d m to drain any accumulated water:

Open wx &el t;mk sun~pdrain and chmk for water and visible cmtanlinants. When tht aux tank is fitred, but has remained unused for prolonged periods, b&m next flight llsiag aux b l , p d ~ r m the following: Carry out water draining procbdure as above.

Fill aux tank with minimum 120 pounds of fuel.

In addition to the normal preflight inspection of the baggage compartment, the pilot should check the f d h cap insib the left baggage c o ~ l p m c ndoor t for fuel level d security. BEFORE STARTING E N G W

Check aux fuel quantity. Chak am h 1 switch seleoted SHUT. 3 E -T A X

Check aux fuel valve position indicator reads SRJT, DyxE Transfer at'auxfuel should begin as soon as cruise has barn established. Select aux fuel valve OPm and check that indicator changes to OPW. A w fuel will transfer to both main fuel tanks at a maximutnrate of 55 pdunds per minute.

Miin &el quantity indicators will rise to near maxinaum and the aux fw1 system wilI maintain . the main ranks at this level until all the aw fie1 has been exhausted.

FAA APPROVED I?EBRUARY 24,1984 Revised January 30, 1993

,

Nornlal Proct.ducs Supplement S- I0

FLIGHT MANUAL

Check that ~UIYfuel valvc is seleotcd SHUT when cngincs are at idle.

FAA APPROVED FEBRUARY 24.1 984

Revised Januaqy 30,1995

SA 4047-76-1

FAA

ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENT

$40 REVISION 1\10, 1

MODEL S-76A PART 2

LONG RANGE FUEL SYSTEM

Date of Issue: FEBRUARY 24,1984

Date of Revision: January 30,1995

-w@@

- .

FLIGHT MANUAL

_ _ -- - . . *

List 9EiG:ctivz Pagzs SilppIcmenr S- 10

PART 2 LIST OF EFFECTIVE PAGES FAA ROTORCWT FLIGHT 'MANUAL

PAGE

nArm

FEBRUARY 24, I 984 Revised J h n w 30.1995

.__ .

REV.NC~,

A FUGHT I ~ / MMANUAL M~~YJS~ LOG OF REVISIONS

Thc rtivision date is shown at the bottom of rtaoh revised page. A vdocil bar on the outer margin indicates the latest revised portion d each page. A vertical bar r i m to rhz page number hdicatzs the page has been rrcstyped with no sign8cant c h g e s .

REVTSION

NO. 1

"

PAGES lXi3VISED

1-1/1-2,2-1.

2-2

-

lUmAFXS Added optional am tank (forward)

installa~ion.

FEBRUARY 24,1984 Revised Jmlrary 30,1995

DATE

Ii30/95

,

'P2u-t.2, Swliorr I Desonption Supplenwnt S-10

SECTTON I

DESCRIPTION

h~s u s i l i q ficl umk with a fuel capaaity crf 106 U.S. gallons is install& in the baggage cornpatmat l'he tmk can bc hstaUedin either the fcnward or aft section of the wmpartrnent, When htaIled inthe forward section all fuel i s usable, when installed in the aft section, them tan ba a m y as four musable gallom, The tank mIsists of a crash-miutautfuel cell installed in a strucrural box.

I

The bm is i s h e d t o a& of composite marerial bolted 10 tlu existing baggage compartment dmr. Ths t;mkmty be quicWy removed form the helicopter, leaving tht;raft installed. All fuel and vent couplings are sclf-sehg quick-disconnects,allowing thr tank to be removed when parriaUy Wl, jf r t i q ~ i d .When ths tank has been removed in this way, the helicopter may be flouv with the remaining am fuel system cornponds still iastplled.

' h e amtank gravityfsed3dirwtly inlo both main tanks via m electric motorized &I valve and o w float-operated shut 08valve installed in each mah tank, The float valves will not permit fuel W a n d e r if the min tanks are fill, They are designed so thz expansion space in the main ttUljiS is always Pee of fuel. '-

Thzmo~orkedvalve isolates the audiwq. tank fiom the main fuel system und is controlled by a cockpit switch. A position hdicam mounted next to tbe switch si@als the position of ?& motorized valve. The valve add indicator ate comeated to the No. 3 dc; p t M q bas ( 8 f t ixsdtition) or the No. 1 dc primary 58 bus (forward iastallatiorr) by a circuit breaker marked

.4uxFUEL VALVE.

A float-type fud probe is iastdled in the auxiliary tank, with r\n indimtor on the pilots b ~ ~pael s andt a repeater next to the filler cap, The fud gage is mwked in &ps Qf 120 pounds, ikom zero to 720 pounds, and indicates the leat quantity of h I rmaiDing (e.g., the indicatorwill change h m 600pounds to 480 pounds as soon u the fiA level drops below 600 pounds). The fid gage is commted to the dc esential bus by $1 circuit breaker marked AUX

FUEL QTY.

FEBRUARY 24,1984 Revised January 30, 1995

-flBFLIGHT QDMANUAL

Piui 1,Suttcm ll Loading hforrnat.i.ion Supplemint S- 10

AUX FIIEr, WIGHT TABLE NOTE :

Whm the adliary tank is insded baggage should nut be carried in tho baggage csmpartmenr.

If tirc~diis placarded indicating tank installation fore or &, use the appropriate loading domation ii this supplement If aircraft

is IlQIp l w d d use AFT INSTALLATION information. Placard is located on h i d e of baggage campartm~ntdoor,

The fornard mounted auxiliary fud installation weights and momcnki arb as follows:

WE1C;m

ITEM Aw Fuel Tank

~OUNQS)

ARM

MOWNT/l 00

82

2302

189.2

19

215.4

WEIGHT (POUNDS)

ARM

MOMENT/100

81.5

234.8 3

191

185

220,OO

41

Asszmbly Empty A w Fuel Sysiern

,

422

Fived Pam AFT INSTAT.LAT1C)N

ITEM. -4ux Fuel T&

Asscmbly Empry Aux Fuel Systenl Fixed Parts

FEBRUARY 24,1984 Revised January 30, 1995

,

a@@FLIGHT MANUAL

When the auxiliiuy talk is installed, particular attention should be paid to the center of gmvigl od~u1ati:ionin order thnt the h-licoptuCG is maintained within the aft lirmt.

NOTE usterivk (*) are based on JP-5 Sud bt 6.8 paunds per TJS, gallon undzr standard canditiom.

'l'he weights n~askedwith

1

-

AUX FUEL, TANK CAPACITY 106 U.S.O -

-

N S -

-

TOTAL

WEIGHT

.. FUEL GAGE 120 240 360

120* 240*

360'

480 600 720 Full (P-4) 730 Fdl (JP-1) 720 Full (P-5)

1

(Poms)

480"

6004 689 716 721*

ARM

MDM3E;NT/l00

273.2 549,l 824.9

227.68 228.78 229.15 229.43 229.53

1101.3 1377,3 1582-4

229.66 229.66 229.60

1644.4 1655.4

AUX FUEL.TANK - CAP .CITY 106 U.S. GALLONS TOTAL WEIGHT

FUEL GAGE

120 240 360 480

-

600 720 FuU (P-4) 720 Full (JF-1) 720 Full (JP-5)

(Pow% 120* 2400 360" 480" 600* 689 716

721*

ARM

MOWNT/I 00

232.29

27 9

233.39 233.76 234.04 234.16 234,27 234.27 234.2 1

560 842

w

FEBRUARY 24,1984 Rwised Janunry 30,1995

1123 1405 1614 1677

1689

,

.

-

SECTION I11

=EL

VALVE MLFTMCTIONS

I

I

The awr fud valve may not operate or may stick, aBkcriag ttansfer o f b l to main tanks.

I

Symptom:

Valve positiw iudicator reads SHUT when switch is selected OPEN.

Action:

QcI6 valve witch. Ifindictitor fa& to show diagonal batching f~llowcdby OPEN, vdvz aclc~atothas failed. Am fuel will not transfer. Continuo flight on mab fuel only,

Symptom;

Vdve positiorl indicator sh~wsdiagonal hatching which does mr chiurge to OPEN afw 5 seomds.

Action:

Valve will be stuck in transit between shut d full open posirion Fuel transfer may occur at reduced rak. Watch main and BUY fuel gages for indication of fuel &&or. Rectify fault in valve b e h e next flight.

I

I

I

I

I

I 11 i

1

FAA AE'PROVEn FEBRUARY 24,1984 Revised J~nusry30,1995

I

.

,

1.

+ @

'pb

A~~E@MsK~~&, FLIGHT MANUAL

*

FAA APPROVED

4.

ROTORCRAFT FLIGHT MANUAL SUPPLEMENT

REG NO. SER NO. This supplement m u s t be attached to the Sikorsky S-76A, FAA approved, RQ~OPcraft Flight Manual, dated November 21, 1978, when a Metair Long Range Fuel System Part No, 33776-61049-013 (Forward) or 33776-61049-014 (Aft) is installed in accordanoe with STC SH 1939 8 0 . The informatbrs contained herein supplements the infomation bf the basic Rotor craft Flight Manuol ,' For

. limitations, procedures, and perforn~ance

., I

\

I

information not contained in this supplement, consult the basic R~torcraStRight Manual.

. .

Approved By :

Date of Approval:

, '. ,;/'

,9'

( ,,,?

I

A& 23W4de4

-

MANAGER, ATLANTAAIRCRAFT CERTIFICATION OFFICE, CENTRAL REGION, FAA

May 23, 1986

FLIGHT MANUAL

,

Kl?M Supplement S16 for Sikorsky S-7fiA SGCTION I OPE RATING LIMITATIONS

TANK INSTALLATJON AND REMOVAL The aux fuel tank shall be installed and removed in accord-

-

ance with Metair Aux Tank No. -013 or 014. Fuel connsctions should be examined for leakage after installation of tank.

WEIGHT AND C .G. LIMITS The rotarcraft when fitted w i t h the auxiliary fuel system is to be operated within the weight and enter of gravity limits specified in Part 1 of the S-,113 Flight Manual. Particular attention should be paid 40 the center of gravity calculntiun before each flight. Refer t o LOADING INFORMATION , Section VI of this supplement, LOADING LIMITS

When the aux fuel system is fitted and aux tank i s in use, a maximum of 150 Ibs of baggqge may be carried if held down to brackets provided an tank assembly, provided the roturwaft C ,C remains w i t h i n specified l i m i t s

.

.

AUX FUEL SYST@VI LIMITS

The aux tank i s only to be g ~ a v i t yfueled. Fueling is only to be carried out with the rotorcraft in a n o r m a l ground attitude t o ensure adequate expansion space in the aux tank.

In-flight t r a n s f e ~of fuel born the aux tank to the main tanks is only to take place with the rotarcraft in normal straight and level flight up t o 140 KXAS.

In the specified flight attitude for fuel transfer there is zero unusable fuel in the aux tank.

FAA APPROVXD

DATE: May

3 3 , 1986

I

Page 2 of 13 ,

.

.

,,

I

.-

FLIGHT MANUAL

RFM Supplement S16 fox Sikorsky LS-76A

I

iI

' .

'

5,

During transfer of fuel f ~ o mthe aux tank to the marin tanks, the aux fuel valve should be left in the OPEN position for at least ten minutes after the aux fuel contents indicator has switched from 60 Ibs t o zero. This is t o msure that dl fuel has transfe~red.

6.

The aux fuel valve must be shut during takeoff and landing t o ensure fuel system independence.

PATE: Mag 23, 1986

Page 3 of 13

FLIGHT MANUAL

RFM. Supplement S 16 for Sikoxsky S-76A SECTION 11

NORMALPRQCEDURES EXTERlOR INSPECTION *.

1.

In addition to the nmmal gre-flight inspection o f the baggage compartment, the pilot should check the fUer cap inside the left baggage compartment door %r fuel level and 8ecuPity.

2.

~ n b the o 4 quarter-turn fasteners on access plate to drain valve in center of aft face of tank. Remove end cap on drain valve, and push hose on end of valve. Drain,

i

!

check for water and visible contaminants. Adequate water drainage i s provided with the helicopter approximately level (less than 3 O nose up or nose down).

This i s t o be done befoxe the first flight of the day or before the next flight after extended maintenance.

BEFORE'

I

i

.

I

STARTXNG ENGINES

Check aux fuel quantity.

i

Check hux fuel switch SKUT.

Check aux fuel valve position indjcator rends SHUT.

Transfey of aux fuel should oommcjnce as soon as cruise has been establtshed, Select aux fuel valve OPEN and cheok that indicator. changes to OPEN, Main fuel quantity indicators will rise to near maximum and the aux fuel system will maintain the main tsnlis at this level until all the wux fuel has been exhausted. The aux fuel valve should be Isft in the OPEN position for at least ten minutes after the aux fuel contents indicator has switched from 60 lbs t o zero, to ensure that all fuel has transferred, L

L.l

RFM Supplement S16 for Siko~skyS-76A

BEFORE LANDING

Before landing, confirm aux fuel valve is a t SHUT. SHUTDOWN Check that atw fuel valve is eeiected SHUT when engines are at idle.

FAA APPROVED DATE: May 23, 1986

Page 6 of 13

,

-'xa

.,

SA 4047,- 76- I ,

FLIGHT MANUAL

RPM Supplement 5 16 for Sikorsky 6-76.A SECTION XI1

\

EMERGENCY PRQCEDURES

AUX FUEL VALVE MALFUNCTIONS I

The aux fuel valve may fail t o operate or stick affecting transfer of fuel t o main tanks. Symptom : Valve position indicator re&s 4s in OPEN position.

Action:

I

,

I

I

SHUT when switch

I1

Cyole valve switch. If indicator fails to show

diagonal hatching followed by OPEN, valve actuator has failed. Aux fuel will not transfer, Continue flight on main fuel only. -

I

iI

I

If indicator changes to diagonal hatahing and doe8 not show OPEN after five seconds valve i s partially open. (See below .)

I

Symptom: Valve position indicator shows diagona1 hatching which does not change to OPEN after five seconds.

Action:

I

Valve will be stuck in transit between shut and open positions. T~ansferof fuel may ~ C C U Pat reduced rate. Watch main and aux fuel gages for indication o f fuel transfer. Rectify fault in valve prior to next flight.

I

FLOAT VALVE MALFUNCTIONS

I

1

I

The float valve may fail t o opepate in any of the following ways :

I

Synrptom: With motmized fuel valve open, the main tank indicators rise above normal fdposition.

!

Cause :

1

If fuel levels oontinue t o rise above normal full level on both m a i n tank indiaatbre with the aux fuel level falling, both float valves would be jammed open.

..

.,

w

i ,

i

FAA APPROVED

DATE: May 23,

1

I

1986

Page 6 of 13

FLIGHT MANUAL

RFM Supplement S 16 for Sikorsky S- 7 6A Close aux motorized fuel vdve until levels of fuel in main tanls drop below normal full level. Continue to ufie this procedure by opening and closing fuel valve to keep main tanks full until aux tank contents are used, or peturn to base to rectify

fault. Ensure fuel levels in m i a h tanks have fallen bebw nmrnsl full level before landing In case of fuel overflow from vents.

Symptom : With motorized valve open, one main tank indicator rises above normal full level and other vdve stays at normal full level. Cause :

If level oonlinues to rise above normal full level on one main tank indicator, one float valve would be jambed open.

Action :

Close aux motorized fuel valve so mziiin tank levels wiSL drop. U s e t h g same procedure as above as if bath valves are jammed,

Symptom : With motorized fuel valve open, one main tank indicator f d s 'below m r m d l full level and other SrtdicaLor stays eonstant at normal full level. Cause ;

If level falls on'on& indicator and the other indicator stays constant at normal ftdl level, one float valve would be jammed closed.

Action ;

Open cross feed valve so that fuel tanks level out, or return 50 base t o have fault rectified.

Symptom : Wit11 motorized fuel valve open, main Lank indicators drop and aux tank indicator stays constant, Cause :

If t h i indiaators show that no fuel has Wansferred from aux tank to either maitin tank, when motorlzed fuel valve is open, both float valves would be jammed dosed.

Action ;

In this case return t o base t o rectify fault,

FAA APPROVBD DATE: May ,23, 1986

Page 7 of 13

PLIGHT MANUAL

RFM Supplement S 1 6 for Sikmsky S- 76A NOTE

If aux motorized fuel valve float valves show signs of malfunctions at the same time, return t o base-immediately to rectify or land as soan as possible, ensuring heUoopter is kept level t o stag any chanoe of fuel over-flowing' from .' vents

.

SECTION IV

PERFORMANCE DATA

NO CHANGE

FAA APPROVBD DATE: May 23, 1986

Page

8

of 13

FLIGHT MANUAL

RFM Supplement S16 foeor Sikorsky S-76A SECTION V

GENEBAL DESCRIPTION

An a ~ u i l i a - yfuel tank, usable oapacity 50 US gdbns, is installed in the baggage compartment. The talk consists of a .&rashresistant fuel cell installed in a structural box, The box is attached to a 'raftt of composite mategal bolted to the existing baggage cornpartmelit floor. The tank itself may be q~$cldyremoved from the ~drcraftleaving the raft installed. All fuel and vent oouplings are self sedifig quk?R itisconnects allowing the tank t o be removed when partially full, if required. When the tank has been removed in this way, the rotarcmft may be.flown with tho remdning aux fuel system eompommts still installed.

,

The tank has a sump situated centrally on the aft face of tank, behind an access plate. The sump has a drain valve.

'

The awuiliary tank feeds diroectly by gravity into both main tanlrs via an eleatpic motorized fud valve and one Boat operated shut off valve fitted in each main tank. The float valves will not permit fuel transfer if the main tanks are full. They are designed such that the exqmnsibn space in the main tanks is always free of fuel. The motorized valve isolates the amdiary tank f r o m the m a i n fuel system and is operated by a switch on the flight deck, A position indicator mounted adjacent to the switch signals the position of the motorized valve. The valve and indioator are connected to the No, 1 DC primary bus by a circuit breakep marked AUX FUEL VALVE. The tank is gfuwity filled through a filler cap an the left side of the tank. Access t o the filler is by opening the left side baggage compartment door. A formed composite soupper in-

side the door prevents spillage during refueling from entering baggage compartment.

: :

'. -

The auxiliary tank is fitted with an electrically signalled contents gage with an indioator on the pilot's instrument and a repeater adjacent to the filler cap. The contents probe consisvtsof a series of magnetic floats. The-cantents gage indicates in step increments of 60 lbs, froom zero to 300 ~ b s wfth , one increment from 300 to 340 lbs. The gage is pcssimfstic indicating the least quantity of fuel remaining. FAA APPROVED DATE: M a S 3 , 1986

-

FLIGHT MANUAL

RFM Supplement S 16 for Sikarsky S- 76A { e . g . , the indicator w i l l change f r o m 300 to 240 Ibs as soon as thecontentsdrop below 300 Ibs). The contents gage is o m nected to the DC essential bus by n circuit breaker marked AUX FUEL QTY.

FAA APPROVED DATE: May 23, 1986

FLIGHT MANUAL

RFM Supplement S 16 for Sikmsky S-76A

2LOADING

INFORMATION FOR THE :013 CFQRWARD) AUX TANK TNsT,ALLATIC,?~-"-/

The auxiliary fuel ltnstallation weights and centroids are as

fqllows :

1 ,Weight ob

Item

Centroid Moment / 100

(STA)

4

Auxiliary Fuel, Tank Assy qmgty

'75.00

225

16 9

;Auxiliary Fuel System

19.50

220

43

L. ,

i

When the auxiliqry tank is 'fitted, a maximum of 300 5bs 031 baggage can be carried in the baggage compartment, provided the aircraft remains within its limftc. I

NOTE

.

C G of baggage is m~vedaft to SSTA 243.3 'when tank I s fitted,

Aux Fuel T&B

-

Fuel Gage

60 120 '

)

180 240 300 340

- Maximum C 2 s c i t y 51 US ~ a l l o n s . Total Weight 60"

225 + 32

139

120"

225.11 225. 00 224.92 224.86

270 405 540 675

224.81 224.. 81

764 771

180* 2 40"

FULL (JP-I)

340 FULL (JP-5) FAA APPROVED DATE: M a y 23, 1986

M~mant/I00

Centroid

300" 540 343

!

Page 11 o f 13

t

FLIGHT MANUAL

RFM Svpglement S1S for Silxmsky S-76A

NOTE

7

C*) The weights marked with an asterisk are based on JP-1 fuel at 6.15 lbs per US

(II

gallon under akndard conditions.

r ,The maximum usable capacity 51 US gallons. 7 ' a

LOADING INFORMATION FOR THE -014 (AFT) AUX TANK INSTALLATION--

Th2 auxiliary fuel ins3allation weights and centroids are as ~ O ~ ~ Q: W S

It em

Weight

Centroid

(lb 3

(STA3

Moment / 100

Auxiliary Fuel Tank hssy Empty

76.00

230

173

Auxiliary Fuel System

19.50

225

44

-i 1. Whei the auxiliary tank i s fitted, a maximum of 150 lbs of baggage can be carriad in the baggage oompartrnent, prodded the aircraft remains w i t h i n its limits.

NOTE C .G. of baggage is moved aft t o STA 246.0

when tank is fitted..

'h.

FAA APPROVED DATE: May 23, 1986-

Page 12 of 13

J%~~B?BGK~~

SA

4047-76-1

FLIGHT MANUAL

RFM Supplement S16 for Sikwsky S-76A AUX.,FUEL ARM TABLE f

."

Aux Fuel Tank

.-

Fuel Gage

- Maxiinurn Cabacity 51' US Gallons Totdl Weight

Centroid

Moment / 100

--

120

60" 120*

180

180"

2 40 300

240"

340 FULL (JP-1) 340 FULL (YP-5)

340 343

60

230,62

230.41 230.30 230.22 230.16 230.11 230.11

SOO*

NOTE 5

(*) The weights marked with an asterisk are based on JP-l fuel at 6.76 lbs per US gallon under standard conditions,

The maximum usable capacity 51 US gallons.

When the auxiliary tank is fitted, particular attention should be paid to the Center of gravity cdcuktion in order that the rotorwaft C .G. is maintained within the aft limit *

FAA APPBOVED DATE: May 23, 1986

Page 13 of 13

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SA 4047-76-1

--

I , I I

I

ROTORCRAFT FLIGHT M A N U A L

SUPPLEMENT NO. 23 SIKORSKY

MODEL PART 2

I

AUXILIARY ESSENTIAL POWER SYSTEM

Date of I s s u e : July 3 , 1984 Date of Revision : March 2 1 , 1985

I FLIGHT MANUAL

Part 2 List of Effective PWes Supplement No. 23

PART 2 LOG OF REVISIONS The revision date is shown at the bottom of each revised page, and a vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO.

1

REVISED PAGE NOS.

1-1, 1-3

REMARKS Revised electrical bus monitored by AUX ESS PWR system. Add List of Effective Pages

July 3, 1984 Revised March 21, 1985

DATE

3/21/85

~ / ~ ~ ~ f l [ d f l & FLIGHT MANUAL

P a r t 2 Section I Description Supplement No. 23

SECTION I DESCRIPTION ELECTRICAL SYSTEM B a t t e r y Switch T h e BATT s w i t c h o n the m a s t e r swiich ( F i g u r e 1-2) h a s positions m a r k e d ON-OFF-AUX ESS TEST. ON c o n n e c t s t h e b a t t e r y to t h e d c e s s e n t i a l b u s except when d c e x t e r n a l power is in use. O F F d i s c o n n e c t s the battery f r o m the d c e s s e n t i a l bus. AUX ESS T E S T c h e c k s t h e integrity of the auxiliary e s s e n t i a l power s y s t e n i . D C POWER DISTRIBUTION SYSTEM

DC power i s distributed through t h r e e b u s e s . T h e Ho. 1 d c p r i m a r y b u s i s powered b y :he No. I g e n e r a t o r u l d the NO. 2 d c p r i m a r y b u s i s powered by the No. 2 g e n e r a t o r . T n e dc e s s e n t i a l b u s i s c o n r ~ e c t e dto No. 1 gcner;ltor output ~ h r o u g ha bus tie c o n t a c t o r . If e i t h e r g e n e r a t o r is inoperative a s indicated by the lighting of the a p l ~ r o p r i a t eDC GEN caution light, the o t h e r g e n e r a t o r will ~ r o v i c i epower to a l l t h r e e b u s e s . U both g e n e r a t o r s a r e inoperative and e.xterna1 p o w e r is not i n u s e , only t h e d c e s s e n t i a l bus will b e powered f r o m the b a t t e r y . T h i s condition i s indicated by the Lighting of both DC GEN and BUS T I E O P E N caution lights. If t h e dr: junction box o r a s s o c i a t e d w i r i n g sustaills d a m a g e that i n t e r r u p t s t h e c l r c u i i r y Lo the dc p r i m a r y b u s e s , the a u x i l i a r y e s s e n t i a l power s y s t e m will be activated Lo connect the b a t t e r y d i r e c t l y to t h e d c e s s e n t i a l bus. T h i s coudition i s indicaied by lighting of t h e AUX ESS PWR ON a d v i s o r y llg11:. E x t e r n a l power :1.111 power al! t h r e e b u s e s u n l e s s o n e o r m o r e g e n e r a t o r s a r e o p e r a t m g : then e x t e r n a l power will b e disconnected f r o m the dc s y s t e m . E s s e n t i a l P o w e r On Advisory Light Ausiliary

The AUX ESS PWR ON a d v i s o r y l i g h ~on t h e caution-advisory ( F i g u r e

( 1-3) lights when power f r o m t h e No. 1 d c p r i m a r y b u s is i n t e r r u p t e d and LIII,

auxiilary power c o ~ i l a c l o rconnects Lhe b a t t e r y to t h e dc e s s e n t i a l b u s .

J u l y 3, 1984 R e v i s e d M a r c h 21, 1965

,

.

GENERATOR ANDGEHERATOR COHTRGL

[T%-%czz==j] CAUTION LIGHT

'L -

CAUTION LIGHT

NO. 1 DC STARTER GENERATOR AND GENERATOR CONTROL

GENERATOR ORS 1. CONNECT EACH GENERATOR TO I T S PRllAARY B U S I F GENERATOR O U T P U T IS ACCEPTABLE TO GENERATOR CONTROL P A N E L 2. CONNECT EITHER GENERATOR TO B O T I I PRItAARY B U S E S I f OTHER GENERATOR IS OFF OR H A S f A l L E D BATTERY OR 1. C O N N E C T S B A T T E R Y TO DC ESSENTIAL B U S ONLY DISCONNECTS BATTERY F R O M D C E S S E N T I A L B U S W H E N EXTERNAL POWER IS I N IJSE OR W I I E N RATTZRY S W I T C H IS OFF.

,

,

..,...- - .,.-pq

, ~ u i , i , L ~ a L>; . ; L , ~ . ' ~ , ' * . < .'-LI< I d 2i : > : d ; C f . ' !.!L, i i O T ! I C;CNERATOfIS nl,r OPERATING. I.!-.< T I E C O N T A C T O I ? . ., , , . L . ., L . ,.,d. ,%..< , , ~ - <

:.

;I

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....

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,

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:

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,

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i ';;.!iLr<

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,

O l < l 0;;

!'~!!~t<

. , , . Q,'::,i

L'OWER IS I N USE. I>,*.,L;,::.S 2;7TiC!.:'i b a T t ' U 7 iI:gP!, P R I ~ ~ I A R ~I l C :;iUSF.S V:liEPJ OPERATING O N B A T T E R Y ONLY.

A U X PWR C O N T A C T O R

1. CONNECTS BATTERY T O D C ESSENTIAL B U S W H E N N O . l DC PRIMARY POWER IS INTERRUPTED.

GENERATOR CALJTION LIGHTS I O N W H E N GENERATOR IS OFF OR INOPERATIVE A N D D C SYSTEM IS P O W E R E D B Y O T H E R GENERATOR. E X T E R N A L POWER.OR B A T T E R Y BATT HOT WARNING LIGHT 1 O N D U R I N G PERIOD OF BATTERY OVERTEMP. W H E N O N B A T T E R Y IS D I S C O N N E C T E D F R O M ESSENTIAL B U S BATT OFF CAUTION LIGHT 1 ': :':IIFf< D C SYSTEM IS Pli'b'.'FRED AND B A T T E R Y S W I T C t I IS OFF.

.

,,

,

-

r

I

'

<.. ' i H i E M P E H A l URE Dl? O i tick

,r(,:::;

r r ; , l r ) ~p c I

---

--

UATTEHY M A L ~ U N L T I O I . I .

1;IJS ;I€ O P E N C A U T I O N L l G l i T I ! ! F GrPiERATOR O U T P U T I i O E S N O T G O ' :1;'3;1k!l 1 I I E 9115 TI!. , , ( ~ : v ; A C T L ~TO ? PO\VER Tr!if L):. F S S i l d T I A L i. ,',r;i,'.lGI:. D C ESL..:;':'i B ! J s is POI'i'ERED li10'.: ti-ITTERY i::;LY.

'.

A U X E S S PWR O N A D V I S O R Y L I G H T

1. O N W H E N NO. 1 D C PRIMARY POWER IS I N T E R R U P T E D A N D BATTERY POWERS DC ESSENTIAL BUS. B A T T E R Y SVIITCH O N OR A T A U X ESS TEST.

Part 2 Section I Descrintion supple&ent NO. 23

RIGHT MANUAL

DUAL INVERTER AC SYSTEM

AC GENERATOR S Y S T E M

Figure 1 - 2 . 1- 4

Master Switch s (Typical) J u l y 3 , 1964

.. .

- r ~

j . , .

-

-.

F l l G t i i MANUAL

Part 2 Section I D:scription Supplement No. 23

-

b "A ->

"0 ,A

I

FAA APPROVED

ROTORCRAFT FLIGHT MANUAL SUPPLEMENT S-24 REVISION NO. 1 SIKORSKY MODEL S-76A PART I HONEYWELL LSZ-850 LIGHTNING SENSOR SYSTEM

Sikorsky A United Technologies Company

Approved by:

Manager, Boston Aircraft Certification Ofice, ANE-150

Date of Approval: October 11, 1988 Date of Revision: July 2, 2003

Part 1 Introduction Supplement S-24

FLIGHT MANUAL INTRODUCTION

This supplement must be attached to the Sikorsky S-76A, FAA approved Rotorcraft Flight Manual dated November 21, 1978, when a Honeywell LSZ-850 Lightning Sensor is installed in accordance with STC SH2416SO or as a 33776-61083 option to the S-76A type design. The information contained herein supplements the information of the basic Rotorcraft Flight Manual. For limitations, procedures, and pelformance information not contained in this supplement, consult the basic Rotorcraft Flight Manual.

I

FAA APPROVED OCTOBER 11, I988 Revised July 2, 2003

FLIGHT MANUAL

Part 1 List of Effective Pages Supplement S-24

PART 1 LIST OF EFFECTIVE PAGES PAGE

DATE

REV. NO.

FAA APPROVED OCTOBER 11,1988 Revised July 2,2003

Part 1 Log of Revisions Supplement S-24

FLIGHT MANUAL PART 1 LOG OF REVISIONS

The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes. REVISION NO.

PAGES REVISED

REMARKS

DATE

1

All

Reissued supplement.

7/2/03

FAA APPROVED

Changed from Atlanta ACO to Boston ACO on title page. Added option to the type design in Introduction.

Manager, Boston Aircraft Certification Office, ANE-150

FAA APPROVED OCTOBER 11,1988 Revised July 2,2003

SA 4047-76-1

Ad~rn~~rnflE FLIGHT MANUAL

Part I Table of Contents Supplement S-24

PART 1 TABLE OF CONTENTS

SECTION

PAGE

I

OPERATING LIMITATIONS.............................................

1-111-2

II

NORMAL PROCEDURES.............................................

2-112-2

Ill

EMERGENCY PROCEDURES.......................................

3-113-2

IV

PERFORMANCE INFORMATION.....................................

4-114-2

FAA APPROVED OCTOBER 11, I988 Revised July 2,2003

ilii

FLIGHT MANUAL

'

supGement S-24

SECTION I OPERATING LIMITATIONS 1.

The Honeywell LSZ-850 Lightning Sensor System Pilot's Handbook, Part No. 28-1 146-5400, dated October 1988, or later approved revision, must be immediately available to the crew at all times.

2.

Flight shall not be predicated on the availability or the use of the LSZ-850 Lightning Sensor System. Verification of displayed lightning has not been demonstrated.

FAA APPROVED OCTOBER 11,1988 Revised July 2,2003

FLIGHT MANUAL

Part 1, Section II Normal Procedures Supplement S-24

SECTION II NORMAL PROCEDURES No change

FAA APPROVED OCTOBER 11,1988 Revised July 2, 2003

FLIGHT MANUAL

Part 1, Section Ill Emeraencv Procedures Supplement S-24

SECTION Ill EMERGENCYPROCEDURES No change

FAA APPROVED OCTOBER 11,1988 Revised

- .

FLIGHT MANUAL

Part 1, Section IV PerformanceInformation -~ - . Supplement S-24

SECTION IV PERFORMANCE INFORMATION No change

FAA APPROVED OCTOBER 11,1988 Revised

FAAAPPROVED

ROTORCRAFT FLIGHT MANUAL SUPPLEMENT NO. 32 REVISION NO. I SIKORSKY MODEL PART 1 DUAL 750 VOLT-AMPERE STATIC INVERTER AC POWER SYSTEM

Sikorsky A United Technologies Company

O SKORSKY AIRCRAFT CORPORATION 1990 Approved by:

Al!6&Gdw'kka;B.

kh

Manager, Boston Aircraft Certification Office, ANE-150

Date of Approval: January 4,1990 Date of Revision: November 6,2003

Part 1 Introduction Supplement No. 32

FLIGHT MANUAL INTRODUCTION

This supplement must be attached to the Sikorsky S-76A, FAA approved, Rotorcraft Flight Manual, dated November 21, 1978, when a Dual 750 Volt-Ampere Static Inverter AC Power System is installed. The information contained herein supplements the information of the basic Rotorcraft Flight Manual. For limitations, procedures, and performance information not contained in this supplement, consult the basic Rotorcraft Flight Manual.

FAA APPROVED JANUARY 4,1990

-7EFLIGHT MANUAL

Part 1 Table of Contents Supplement No. 32

PART 1 TABLE OF CONTENTS

SECTION

TITLE

PAGE

I II 111

OPERATING LIMITATIONS .......................................... NORMAL PROCEDURES ............................................. EMERGENCY PROCEDURES......................................

1-1 2-1 31

FAA APPROVED JANUARY 4,1990 ilii

J'm mm'E FLIGHT MANUAL

Part 1 List of Effective Pages Supplement No. 32

PART 1 LIST OF EFFECTIVE PAGES PAGE

DATE

REV. NO.

A B ilii 1-1 2-1 3-1

11/6/03 11/6/03 1/4/90 11/6/03 1116/03 1/4/90

1 1 Basic 1 1 Basic

FAA APPROVED JANUARY 4.1990 Revised November 6,2003

Part 1 Log of Revisions Supplement No. 32

FLIGHT MANUAL

PART 1 LOG OF REVISIONS The revision' date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes. REVISION NO. 1

PAGES REVISED 1-1,2-1

REMARKS Revised Operating imitation^ and Dual Static Inverter check.

DATE 11/6/03

FAA APPROVED

Manager, Boston id raft Certification Office, ANE-150

FAA APPROVED FEBRUARY 26,1996 Revised November 6,2003

FLIGHT MANUAL

Part 1, Section I O~eratinaLimitations 'Suppl&nent No. 32

SECTION I OPERATING LIMITATIONS

Day, Night IFR. The following equipment must be installed and operating for operation under Instrument Flight Rules (IFR). 3. Two independent sources of AC power, and Inverter Switching operational.

FAA APPROVED JANUARY 4,1990 Revised November 6,2003

I I

FLIGHT MANUAL

Part 1, Section II Normal Procedures Supplement No. 32

SECTION II NORMAL PROCEDURES

-

+25. AC Generator Check at 100% NR.

.

b. AC generator should drop off the line as indicated by the lighting of the AC GEN caution light.

i-25A.

Dual Static Inverter- Check.

NOTE During steps b, d, and f, a 1 INV FAIL, 2 INV FAIL, or INVERTER will appear on the cautionladvisory .

-

a. AC generator OFF. b. No. 2 inverter ON and No. 1 inverter - OFF. Tune ADF to a local frequency. Observe that ADF pointers on the Pilot and Copilot's HSI point in the general direction of the station. On aircraft with Remote Instrument Controller, observe that when rotating the Course #1 and Course #2 knobs on the Remote Instrument Controller, the Pilot and Co-pilot's course pointers on each EHSl will move in the direction of knob movement. c.

No. 1 inverter- ON

AC generator - ON (if ac generator installed). Both inverters - OFF, check as in step b. above (if ac generator in-

d. No. 2 inverter OFF, check as in step b. above. e. f.

stalled).

g. AC generator, No. 1 inverter, and No. 2 inverter - ON (if ac generator in-

stalled).

NOTE When both course pointers or ADF pointers are inoperative during steps b or d, an inverter is inoperative. When one course pointer or ADF pointer is inoperative during steps b or d, Inverter Switching is inoperative.

FAA APPROVED JANUARY 4,1990 Revised November 6,2003

I

FAAAPPROVED

ROTORCRAFT FLIGHT MANUAL SUPPLEMENT REVISION NO. 4

SIKORSKY MODEL PART 1 SPZ-7600 DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM (DAFCS)

Sikorsky A United Technologies Company

O SIKORSKY AIRCRAFT CORPORATION 1990

Approved by

Date of Approval: OCTOBER 16,1990 Date of Revision: July 15,2004

Manager, Aircraft Certification Office Federal Aviation istration Atlanta, Georgia

-flB

FLIGHT MANUAL

Part 1 List of Effective Paaes ~"pplernents.38

PART 1 LIST OF EFFECTIVE PAGES

PAGE

DATE

REV. NO. 4

3 4

Basic Basic Basic 4 Basic Basic Basic Basic 1 Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic Basic

FAA APPROVED OCTOBER 16,1990 Revised July 15, 2004

Part 1 Log of Revisions Supplement S-38 PART 1 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes. REVISION NO. 1

PAGES REVISED 2-3

REMARKS Revised GoAround description to match pilot's handbook.

DATE

FAA APPROVED

6/4/92

Manager, Aircraft Certification Office, Federal Aviation istration Atlanta. Georgia 2

1-3

Revised Minimum Flight Crew Operating Limitations

8/6/99

Manager. Boston Aircraft Certification Office. AN€-150 Revised required equipment for single Pilot IFR operations

11/6/03

Manager, Boston ~ i & a f t certification Office, AN€-1 50

FAA APPROVED OCTOBER 16,1990 Revised November 6.2003

Part I Loa of Revisions supplement S-38

FLIGHT MANUAL LOG OF REVISIONS

The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes. REVISION NO.

PAGES REVISED

REMARKS equipment ;or dual Pilot IFR operations

DATE

FAA APPROVED

Manager. Boston Ai raft Certification Office, ANE-150

FAA APPROVED OCTOBER 16,1990 Revised July 15, 2004

-BE

FLIGHT MANUAL

Part 1, Section I operating Limitations supplement S-38

SECTION I OPERATING LIMITATIONS Limitations specified in the Sikorsky S-76A Flight Manual and Supplements in use remain applicable. Operation of the SPZ-7600 system does not relieve the flight crew of the responsibility to monitor and comply with all aircraft and engine limitations. The RFM and supplement limitations are complemented by the following limitations:

SPZ-7600 operation is approved to the S-76A Arriel maximum takeoff and landing weight of 10.800 pounds (4,763 kilograms). Automatic approach and hover operations are limited to out-of-ground effect (OGE) hover weight or less. Refer to Section I of the Sikorsky S-76A Flight Manual and Sikorsky Rotorcraft Flight Manual Supplement No. 29 for further information. CG CENTER OF GRAVITY LIMITS SPZ-7600 operation is approved within the standard S-76A Arriel CG limits. Refer to Section I of the Sikorsky S-76A Flight Manual for further information. TYPFS OF OPFRATION CATEGORY 'A" AND "B" Transport Land Day. Night. IFR, VFR IFR Operations Not Approved With: 1.

External Cargo Loads (unless using SAR modes)

2.

Any Doors Open (unless using SAR modes)

-

VFR - or IFR One pilot in right-hand seat IFR automatic approach and hover operations - Two pilots

FAA APPROVED OCTOBER 16,1990

Part 1, Section I O~eratinaLimitations s;lpplem&t S-38

FLIGHT MANUAL

To be eligible for operations under IFR, the following equipment must be installed and operating, in addition to any other instrument or equipment which may be required by Federal Aviation Regulations: 1.

2.

Single Pilot (IFR).

-

Two autopilot systems operating in the ATT Mode

-

Two independent sources of AC power, with Inverter Switching operational

-

Cyclic stick force trim

-

Standby Attitude Indicator with emergency battery Standby CDI (or RMI) located on the pilot's side of the instrument

-

Two DC generators

-

A navigation and communication system that has demonstrated compliance with the pertinent FAA requirements

-

Emergency ICS or transmit switch mounted on the pilot's side and spare headset

Two Pilots (IFR). All of the previous items apply except as follows:

I

-

Either autopilot system must be operable in the ATT mode

-

Standby CDI (or RMI) is not required

-

Copilot's instrument and controls must be installed and operable Emergency ICS or transmit switch mounted on the pilot's side and spare headset is not required NOTE

Autopilot operation not mandatory for VFR operations. Refer to Section I of the Sikorsky S-76A Flight Manual for further information.

3.

-

Automatic Approach and Hover Operations Two Pilots (IFR). In addition to normal two pilot IFR equipment:

Two autopilot systems operating in the ATT mode

FAA APPROVED OCTOBER 16,1990 Revised July 15, 2004

FAAAPPROVED

ROTORCRAFT FLIGHT MANUAL SUPPLEMENT S-52 REVISION NO. I SIKORSKY MODEL S-76A PART I HONEYWELL DATA NAV Ill

Sikorsky A United Technologies Company

u

Approved by:

Manager, Boston Aircraft Certification Office, ANE-150

Date of Approval: october 11,1990 Date of Revision: July 2, 2003

Part 1 Introduction Supplement S-52

FLIGHT MANUAL INTRODUCTION

This supplement must be attached to the Sikorsky S-76A, FAA approved Rotorcraft Flight Manual dated November 21, 1978, when a Honeywell Data Nav Ill system is installed in accordance with STC SH2847SO or as a 33776-61 134 option to the S-76A type design. The information contained herein supplements the information of the basic Rotorcraft Flight Manual. For limitations, procedures, and performance information not contained in this supplement, consult the basic Rotorcraft Flight Manual.

I

FAA APPROVED OCTOBER 11,1990 Revised July 2,2003

FLIGHT MANUAL

Part 1 List of Effective Paaes Supplement S-52

PART 1 LIST OF EFFECTIVE PAGES PAGE

DATE

REV. NO.

FAA APPROVED OCTOBER 11,1990 Revised July 2, 2003

Part 1 Log of Revisions Supplement S-52

PART 1 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. A vertical bar next to the page number indicates the page has been retyped with no significant changes. REVISION NO.

PAGES REVISED

REMARKS

DATE

1

All

Reissued supplement.

7/2/03

FAA APPROVED

Changed from Atlanta ACO to Boston ACO on title page. Added option to the type design in Introduction.

Manager, Boston Aircratl Certification Office, ANE-150

FAA APPROVED OCTOBER 11,1990 Revised July 2, 2003

Part 1 Table of Contents Supplement S-52

FLIGHT MANUAL

PART 1 TABLE OF CONTENTS

SECTION

PAGE

I

OPERATING LIMITATIONS.............................................

1-111-2

II

NORMAL PROCEDURES.............................. .. .

2-112-2

111

EMERGENCY PROCEDURES........................................

IV

PERFORMANCE INFORMATION............. .....................

.

FAA APPROVED OCTOBER 11,1990 Revised July 2,2003

. ...

3-113-2 4-114-2

AHK@BN~~~E FLIGHT MANUAL

Part 1, Section I Operating Limitations Supplement S-52

SECTION I OPERATING LIMITATIONS No change

FAA APPROVED OCTOBER 11,1990 Revised July 2,2003

Part 1, Section Afl~rn~~rnflE Normal Procedures

II

FLIGHT MANUAL

Supplement S-52

SECTION I1 NORMAL PROCEDURES The Data Nav Ill controller shall select the -programmable checklist (Norm & EMER), weather, or the long range nav data to the radar indicator. Refer to the Data Nav Ill pilot's operating manual for operation instructions.

FAA APPROVED OCTOBER 11,1990 Revised July 2, 2003

FLIGHT MANUAL

Part I, Section Ill Emeraencv " .Procedures Supplement S-52

SECTION Ill EMERGENCY PROCEDURES No change

FAA APPROVED OCTOBER 11,1990 Revised July 2,2003

SA 4047-76-1

A FLIGHT ~ ~ rMANUAL n~~rnflE

Part 1, Section IV PeMmance Information Supplement S-52

SECTION IV PERFORMANCE LIMITATIONS No change

FAA APPROVED OCTOBER 11,1990 Revised July 2,2003

SA 4047-76-1 FAA

ROTORCRAFT FLIGHT MANUAL REVISION NO. 29

SIKORSKY MODEL S-76A PART 2

SYSTEMS DESCRIPTION, LOADING INFORMATION, NOISE ABATEMENT, AND CRUISE PERFORMANCE DATA

Sikorsky A United Technologies Company

63 Sikorsky Aircraft Corporation, 1978

Date of Issue: NOVEMBER 21,1978 Date of Revision: November 4,2002

h FLIGHT f f ? MANUAL ~mflE

Part 2 List of Effective Pages

PART 2 LIST OF EFFECTIVE PAGES PAGE A 6 C D E F G H J

K L M NIP I

ii iii

iv v

vi vii viii viiiA ix X

xi xii 1-1 1-2 1-2A 1-26 1-2C 1-2D 1-2U1-2F 1-3 1-4 1-5 1-6 1-6A 1-7 1-8

DATE 11/4/02 17/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 2/22/85 2/22\85 6/29/84 7/26/82 7/26/82 7/26/82 Deleted 11/4/02 11/4/02 3/2/89 3/2/89 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 11/4/02 6/28/84 7/26/82 7/26/82 6/28/84 7/26/82 7/26/82 2/21/7 9

NOVEMBER 21,1978 Revised November 4,2002

REV. NO.

Part 2 List of Effective Pages

3 FLIGHT ~ ~ MANUAL m~~rnflE PART 2

LIST OF EFFECTIVE PAGES (Cont.) PAGE

DATE

REV. NO.

5115/92 1117/83 5115/92 6/28/84 6/28/84 2/22/85 11/21/78 11/21/78 9/7/90 6/28/84 8/27/79 7/26/82 5/ 15/92 8/27/79 5115/92 2/22/85 2/22/85 2/22/85 3/2/89 3/2/89 9/7/90 7/26/82 7/26/82 5115/92 5115/92 2/22/85 5115/92 5115/92 7/26/82 7/26/82 1/17/83 5115/92 6/28/84 2/22/85 Deleted Deleted 2/1/79 7110187 3117/86

28 18 28 20 20 22 Basic Basic 27 20 7 16 28 7 28 22 22 22 26 26 27 16 16 28 28 22 28 28 16 16 18 28 20 22 22 22 3 25 24

NOVEMBER 21, 1978 Revised November 4,2002

Part 2

k ' ! 8 f f ~ ~ ~ f List l of~ Effective Pages FLIGHT MANUAL PART 2 LIST OF EFFECTIVE PAGES (Cont.) PAGE

DATE

REV. NO. 20 21 21 20 13 13 13 20 16 16 13 13 Basic Basic 5 27 25 25 Basic 25 Basic 28 28 11 16 16 20 18 20 18 18 16 16 6 15 11 16 16 10

NOVEMBER 21, 1978 Revised November 4,2002

w

FLIGHT MANUAL PART 2 LIST OF EFFECTIVE PAGES (Cont.) PAGE

DATE -

REV. NO. 10 10 10 23 23 23 Basic Basic Basic 10 10 26 26 26 26 26 26 26 26 26 26 26 26 26 26 26 26 26 10 10 10 10 Basic Basic Basic 1 Basic 16 Basic

NOVEMBER 21,1978 Revised November 4,2002

h ff!@!~~ FLIGHT MANUAL

SA 4047-76-1

Part 2 List of Effective Pages

PART 2 LIST OF EFFECTIVE PAGES (Cont.) PAGE

DATE

REV. NO.

2-36 2-36A 2-37 3-1 3-2 3-3 3-4 3-5 3-6 3-7 3-8 3-9 3-10 3-11 3-12 3-13 3-14 3-15 3-16 3-17 3-18 3-19 3-20 3-21 3-22 3-23 3-24 3-25 3-26 3-27 3-28 3-29 3-30 3-31 3-32 3-33 3-34 3-35 3-36

4112/79 6/6/80 119/79 6/6/80 6/6/80 10114/80 5114/85 5114/85 5/14/85 5114/85 5114/85 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 616180 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80 6/6/80

6 10 1 10 10 11 23 23 23 23 23 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10 10

NOVEMBER 21,1978 Revised November 4,2002

-

FLIGHT MANUAL PART 2 LIST OF EFFECTIVE PAGES (Cont.)

NOVEMBER 21, 1978 Revised November 4,2002

Part 2 Log of Revisions

FLIGHT MANUAL PART 2 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO.

PAGES REVISED

REMARKS

DATE

1

1-2, 1-56 1-56A, 1-57, 2-2, 2-3.2-4,2-5,2-8, 2-11,2-17,2-32,2-37

Added Flotation System and miscellaneous changes.

1/9/79

2

1-12A

Added fuel flowmeter.

1/17/79

3

Intro, 1-2A, 1-24, 1-36, 1-37, 1-49

Added metric values, Inverter failure waming light.

211/79

4

i, vii, 1-8, 1-9, 1-10, 1-11, 1-12A

Added fuel totalizer/flow meter.

2/21/79

5

viii, 1-2A, 1-24, 1-48, 1-56A, 1-58

Added metric values and CU.

3/16/79

6

vi, viii, 1-1, 1-2, 1-2A, 12B, 1-55, 1-60, l-61,2-36

Added cargo hook and sliding door.

4/12/79

--

7

-

i, vii, I-2A,1-16A, I-16B,

-

~-

Added EAPS.

8/27/79

1-16C, 1-17 -

9

i, ii, 1-16C, 142

-

- -

-

-- -

Deleted servo jam warning system.

12/20/79

Added snow protection kit PIN 76090-30008

5/22/80

NOVEMBER 21,1978 Revised November 4,2002

Part 2 Loa " of Revisions -

-

-

FLIGHT MANUAL PART 2 LOG OF REVISIONS

The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO.

PAGES REVISED

.

REMARKS

DATE

I0

v v iv ix I 1-62, 1-63, 1-64, 1-65, 1-66, 2-1,22,2-2A, 2-8,2-9, 2-27,2-28, 2-28A, 2-28B, 2-36A, 3-1, 3-2,3-3.34.3-5. 3-6, 3-7.3-8, 3-9. 3-10,3-11, 3-12, 3-13, 3-14, 3-15, 3-16,3-17-3-18, 3-19, 3-20, 3-21, 3-22, 3-23, 3-24, 3-25, 3-26.3-27. 3-28.3-29,3-30, 3-31, 3-32, 3-33, 3-34, 3-35, 3-36, 3-37, 3-38,3-39, 3-40, 3-41, 3-42, 3-43.344.3-45, 3-46, 347,348, 3-49

Added hoist.

11

i, iv, 1-1, 1-2, I-2A,

Added left sliding doorElectric door locks collective to yaw coupling - ECU operation and miscellaneous changes.

10114/80

Flight control servo low pressure caution lights description modified.

12117/80

1-2B, 140, 1-41. 1-50, 1-53, 1-54, 1-56B, 1-62,l-63, 1-64,3-3,340, 348,349 12

1-21, 140A. 1-41

NOVEMBER 21, I978 Revised November 4,2002

FLIGHT MANUAL

Part 2 Log of Revisions

PART 2 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO.

PAGES REVISED

13

iv, viii, ix, 1-42, 1-43, 1-44, 1-44A, 1-446,l-44C. 1-440, 1-44E, 1-45

Added AFCS Phase Ill and Editorial changes.

3/24/81

14

2-3, 2-4, 2-5, 3-4, 3-5, 3-6, 3-7, 3-8

Increased maximum gross weight to 10, 300 pounds.

11/2/81

15

1-6, 1-16C, 1-16D, 1-53. 1-56A, 1-62

Added Snow Protection Kit with Continuous Ignition, Hoist Lockout of ECU and Vent Blower. Auxiliary CB

11/16/81

REMARKS

NOVEMBER 21,1978 Revised November 4,2002

DATE

Part 2 Log of Revisions

A//rn~~rnflE FLIGHT MANUAL PART 2 LOG OF REVISIONS

The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO. 16

PAGES REVISED

i. ii, iii, iv, v, vi, vii, viii. ix, x, xi, 1-2, 1-2A, 1-2B, 1-2C, 1-4, 1-5, 1-6A, 1-7, 1-12, 1-16, 1-16C, 1-16D, 1-19, 1-20, 1-22, 1-23, 1-24, 1-25, 1-26, 1-27, 1-28, 1-29, 1-30, 1-38, 1-39, 1-40, 1-44B, 1-44C, 1-49, 1-54A, 1-55, 1-56, 1-56A, 1-59, 1-60, 1-64, 1-65, 2-21, 2-34 REMARKS Added Acceleration Bleed Valve Emergency Lighting Droop Compensator Amplifier Description DC Generator Hot Caution Lights Revise Table of Contents List of Illustrations Sliding Doors Collective Pitch Lever Cyclic Control Stick Trim Flotation System Heating and Ventilating Rotor Brake Switch Pressure Figures Center Console Collective Stick Grip Fuel System All Circuit Breaker s DC Electrical System AFCS Phase Ill Control Utility Hoist Engine Oil and Windshield Washer Fluid Weight CautionIAdvisory Pilot's Compartment Master Switch Fuel Gage Test and Ni/Ts Digital

NOVEMBER 21,1978 Revised November 4,2002

DATE 7/26/82

Part 2 Log of Revisions

FLIGHT MANUAL

PART 2 LOG-OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. --

REVISION NO.

-

PAGES REVISED

REMARKS

DATE

17

viiiA, ix, 4-1, 4-2

Added SECION IV on Noise Abatement.

9/27/82

18

ix, xi, 1-2, I-2A, 1-10, 1-31, 1-56, 1-56A, 1-568, 1-57, 1-58, 3-44

Added RHDR light, 250-C30S engine, and Cabin Ventilating System. Revised ECU control and Windshield Washer Fluid Table. Added nitrogen substitute for helium in flotation system.

1/ I 7/83

19

5-1, 5-2, 5-3, 5-4, 5-5, 5-6, 5-7, 5-8, 5-9, 5-10, 5-11, 5-12, 5-13, 5-14, 5-15, 5-16, 5-17, 5-18, 5-19, 5-20, 5-21, 5-22

Added SECTION V on cruise performance data with fuel requirements.

3130183

20

1-1, 1-2A, 1-28, 1-3, 1-6, 1-12, 1-12A, 1-16, 1-16A, 1-19, 1-33, 1-34, 1-34A, 1-39, 1-41, I-44A. 1-56. I-56B

Revised engine description, T5 indicators, rotor brake switch, DC power system, PBA, flotation system, ECU, and editorial corrections.

6/28/84

21

iv, v, 1-40, 1-40A

Added self-centering tail rotor control quadrant.

6/29/84

--

-

NOVEMBER 21,1978 Revised November 4,2002

Part 2 Log of Revisions

FLIGHT MANUAL PART 2 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO.

PAGES REVISED

22

i iii v I , x I 1-13, 1-16, 1-19, 1-20, I-20A. 1-208, 1-21. 1-25, 1-26, 1-26A, 1-27, 1-28, 1-34, 1-49

Added manual rotor brake. Revised engine chip detectors and fuel system. Deleted dual inverter AC system.

2/22/85

23

2-3,2-4,2-5, 3 4 , 3-5, 3-6, 3-7, 3-8

Increased maximum gross weight to 10,500 pounds.

5114/85

24

1-2A, 1-38, 1-49, 1-50, 1-53

Revised collective trim description and master warning .

3117/86

25

1-21, 1-26, 1-37, 1-50, 1-50A, 1-52

Revised fuel system, circuit breaker , inverter fail light, cautionadvisory and OAT gage light.

7110187

26

i x x x i x i 1-16, 1-20B, 1-21, 2-10, 2-11, 2-12, 2-13, 2-14, 2-15, 2-16, 2-17, 2-18, 2-19, 2-20, 2-21, 2-22, 2-23, 2-24, 2-25, 2-26

Updated master switch , fuel system, and chart A.

3/2/89

27

1-16, 1-22, 1-49

Revised fuel system and caution/advisory figures.

9/7/90

REMARKS

NOVEMBER 21, I978 Revised November 4,2002

DATE

AIMBS~M~: FLIGHT MANUAL

Part 2 Log of Revisions

PART 2 LOG OF REVISIONS The revision date is shown at the bottom of each revised page. A vertical bar on the outer margin indicates the latest revised portion of each page. REVISION NO.

PAGES REVISED

28

ii, 1-2B, 1-9, 1-11, 1-16D, 1-18, 1-25. 1-26, 1-27, 1-28, 1-32. 1-52B. 1-53

Deleted electronic N2 overspeed.

5115/92

29

i, ii, ix, x, 1-1, 1-2, 1-2A, 1-2B. 1-2C. 1-2D, 1-2El1-2F, 4-1, 4-2

Updated Table of Contents and List of Illustrations. Added Guidelines for enger Boarding and Exit, Cargo Loading and Ground Operations with Engines andlor Rotors Turning. Revised Noise Abatement procedures.

11/4/02

REMARKS

NOVEMBER 21,1978 Revised November 4.2002

DATE

NIP

FLIGHT MANUAL PART 2 TABLE OF CONTENTS PAGE LIST OF EFFECTIVE PAGES............................................................ LOG OF REVISIONS ........................................................................

A G

SECTION I

DESCRIPTION .......................................................................

1-1

GENERAL DESCRIPTION.......................................................

1-1 1-1

GUIDELINES FOR ENGER BOARDING AND EXIT. CARGO LOADING AND GROUND OPERATIONS WITH ENGINES AND/OR ROTORS TURNING .......................... HINGED DOORS................................................................... Electric Door Locks ................................................................ SLIDING DOORS ..................................................................

1-1 1-28 1-26 1-2C 1-2C

ENGINE RATINGS AND RECOMMENDED USAGE............ ENGINE FUEL SYSTEM ....................................................... ACCELERATION BLEED VALVE ......................................... ENGINE OIL SYSTEM........................................................... ENGINE CONTROLS ............................................................ Enqine Levers .................................................................. N2 Speed Trim Switches ................................................... ENGINE INDICATORS.......................................................... Torauemeter ......................... ........ ............................. T5 (Power Turbine Inlet Tem~erature)Indicators ............. Diaital T5 Repeater ........................................................... NI (Gas Producer) Tachometer ........................................ Triple Tachometer............................................................. Fuel Pressure Indicators ..................... . . . ....................... Fuel Flow/lotalizer System............................................... Fuel Pressure Warnins Liahts .......................................... Enaine Oil Pressurerrem~eraturelndicators ................... Enqine Oil Low Pressure Warnina Liaht ........................... Enqine Chip Detector Caution Liahts................................ Enaine Bleed-Air Advisorv Liaht ....................................... ENGINE-STARTER AND IGNITION SYSTEM ..................... NOVEMBER 21. 1978 Revised November 4.2002

1-2D 1-20 1-2E 1-2E 1-6 1-6 1-6 1-6A 1-6A 1-12 1-12 1-12 1-12 1-12A 1-1% 1-13 1-13 1-13 1-13 1-14 1-14

I

Part 2 Table of Contents

FLIGHT MANUAL

TABLE OF CONTENTS (Cont) SECTION

PAGE

Master Start Switch ........................................................... Enaine Starter Buttons ...................................................... ENGINE ANTI-ICE SYSTEM ................................................. Enaine Anti-Ice Switches ............................................... Anti-Ice Caution Liahts................................................... Anti-Ice On Advisory Liahts ........................................... ENGINE AIR PARTICLE SEPARATORS (EAPS) ................. EAPS Switch.................................................................. EAPS Caution Liahts ................................................... EAPS Advisorv Liaht................................................... SNOW PROTECTION KIT .................................................... ENGINE OVERSPEED PROTECTION SYSTEM ................. TRANSMISSION SYSTEM....................................................... TRANSMISSION OIL SYSTEM ............................................. Main Gear Box Oil PressurelTem~eratureGaae ............. Main Gear Box Oil Low Pressure Warnina Liaht .............. Main Gear Box Oil Hot Warnina Liaht............................ GEAR BOX CHIP DETECTOWOIL HOT SYSTEM .............. ELECTRIC ROTOR BRAKE SYSTEM .................................. Rotor Brake Switch ........................................................... Rotor Brake Caution Liaht................................................. MANUAL ROTOR BRAKE SYSTEM..................................... Master Cvlinder ................................................................. Accumulator ...................................................................... Rotor Brake ....................................................................... Rotor Brake Caution Liaht................................................. POTOR SYSTEM ...................................................................... MAIN ROTOR SYSTEM ........................................................ Main Rotor Blades ............................................................ TAIL ROTOR SYSTEM..........................................................

NOVEMBER 21. 1978 Revised November 4. 2002

Part2 Table of Cantents

FLIGHT MANUAL

TABLE OF CONTENTS (Cant) PAGE

F-mL LEVERS ----.,.- ENGINE PRLME m C H E FUEL QUANTITY DiDICATOR..

... F'uel Gage Test Button .........

FUEL LOW LIGHTS

LEVEL CAUTION

,.,...........,......,. EIZCI'RTCAL - SYSTEM ............. m C U r r BIZEAKEEt S.. ,... DC POWER SUPPLY SPSTEM.. -. . No. 1 and No. 2 St;uter

......,........... Generator Switches ......... G e n W Test Switches,. .

Generators

,,

Generator Carrtion Lights,. DC Generator Hot Ligfrts

-

..

...................

.. ,.....

3 a t b - y SWifEh.. ,, ,, BatteryHotWarniug Light...

.. Power ...............

Battery OLf Caatioo Light..

NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Part 2 Table d Cantents

F U M MANUAL

PAGE

SECTION AC POWER EXIFPLY m ERATOR SPSTEM) ..

M (AC GEN-

. ..... ......... ,

....... ...... ......

.

AC Genemtor Switch .,., AC Generator Test Switch.. AC Generator Carrtion Light AC Generator Bearing Qution Light......,..................

........................., Inverter Switch ................. Inverter Fail Light .............. Invertex On Advisory Light.. .....

Inverter

m D-ON EELATOR SYSTEM)

AC P

(AC GEN-

. ....,........... ,

.. ............. MAIN ROTOR FLIGHT CONTROIS ..... Collective Pitch Lever ............

FLIGHT CONTROL

,

,

Collective Pitch Lever 'Prim a d Force Gradient S p r i n g . . ...- ...

.

Wclic Conk01 Stick Trim and Force Gradient 5,riPg.. .,

.- ..., Pitch Bias AetrJator ...,.... ....... ClPlznneiMnnitarswiteh ........., ,

Tail Botoz Pedals

..... . ...............

T

t

TALROTORXUGET CONTI#)LS a

i

l

~

e

n

,

g

n

o

b

s

.

COLIZCI?VETOYAW COUPIJlG..... SELF-cE24-G TAIL ROTOR COrJTRoL QUAI)-

..........- .-

NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

.

Part 2 Table of Contents

NGHT MANUAL

TABLE OF CONTENTS (Cant) SE,CTION

PAGE

-....,.,,--.-.........-. .........

FLIGHT COhTROL ,HHIRAUI+ICSWVO

..

SYsI'EM

1-40A

FUGET CONTROL SERVO SWITCH-. - - - . .. .. -FLIGHT COh-L SERVO HYDRAULIC PRESSURE lNDIC!ATOR ..,....... FLIGHT COhTROL SERVO LOW PRESSURE CAUTIOZI;LIGHTS .....,....,.. -

~

-

-

~

-

-

-

-

~

-

AUTOMATIC FLIG3T CONTROL SYSTEM

CS P P . 1-41 n

............. ...................,... ......... ... ............. ............

AT'LTTITDE RETENTION.. YAWCHANNEL AFCS CONTROL .. AFCS INDICATOR AFCSCIRCUrrBREAKERS

,

1-43 1-43 1-43

144 1-44

.......... .....

144D

LANDING GEAR CONTaOL HANDLE AND INDICATOR LIGHTS , . . .. LANDING GEAR UP WARMNG UGHT.. EMERGENCY LANDING GEAR SYSTEM.

1-45 145

LANDING GEAR sxs'nnd

- 0 . .

. .. . ...

.

............................ PrruT STATIC SYSTEM.. ........- .... Pftot Heaters ......................

mRUMENTS

NOVEMBER 21, 1978 Revised JUNE 29, 1984

1-46 1-46

1-46

1-46

Part 2 Table of Contents

FLIGHT MANUAL

PAGE

........ STANDBY COM ..,,... .......... FREE-AIR TEMPERA= GAGE ..... TAIL CONE STATIC STSTEM.. ,

. ........... .. ................ ............ C-14A COM SYSTEM (OPTIONAL). Atti-

Indicators. ,

,

Directional Gyms.. Turn and Slip Indicators

ATTITmlE lMllCATOR (5") 561476 1053 (OPTIONAL)

..................

COURSE DEVIATION INDlCATOR (OPTIONAL) . a . . . . .

RADIO MAGmnc IMlICATOR (OPTIONAL).... ..- - ,,..

- ... ........ WHEEL BRAgE SYSTEM ................... CAUTION-ADVISORY ... ........ .. MASlcER WARNING .. ...-. .......... ,

,

.............. PositionLights .-........,......... Anti-collisim Lights ... ........... Cantrallable SeamUght ............

ExrEmOR LIGHTS..

0..

,

........ ......

Se;ucblight M2ster ma. Seanumlt control Switch..

..

mInstsmnent Ligm.. . N o n - U t IustTument Lights.. Lower Console Lights . ,

...

..,... .. . ....,.... - 0 . .

Part 2 Table of Contents NGHT MANUAL

TABLE OF COmENTS ( C o d

SE (TION

PAGE

ox,

..................

..,,.......... B a g g e Compartment Lipht ........., EMERGENCY EQUIPMEET AND SYSTEMS ..... ENGINE DETECTOR S l S I Z M ...... F i e Detector Test Switch ......... . ... ENFIRE EXTINGmSBER SYSTEM.. . Engine T-Handles .................... : Utility hght..

Cabin Domt

-

,

1-53 1-53 1-53 1-53 1-54

-............ 1-54

FireExUquisher Wikh., Fire -her Test h i t c h and Indicator Ligws

.....................

1-54

Flabtion S y s k n Arming Switch aad Flo&tionSystem Test Switch and

....

Inr3icatorLigWs

r...........,

HEATER CONTROL - m A L

..,..........1-56A

CONTROL UMT (ECU).,....

1-56A

...,,..,......... 1-56A .........,,,.....b56B

ECU CONTROL E CU OPERATION ....,

NOVEMBER 21, 1978 Revised July 26, 1982

Part2 Table of Contents

FLIGHT MANUAL

TABLE OF CONTENTS (Cont) \

PAGE

SECTION

mSBIELDWIPER

.....,........-. ...;......... .............

1-58 1-59 1-59

......................... .

1-60

m S B I E L D WASEEEL.. WINDSHIELD HEATER..

CARGO ROOK

,

CARGO HOOK ARMING SWITCH AND ADVISORYLIGHT CARGO SWrrCfi AND HOQK OPEN ADVISORY LIGHT. , .. GROUND RELEASE I;EVER - ,,, -, - , - ... EMERGEN& CARGO HOM( RELEASE T-IIAM)LE

....... .........-.

....... .... ... . .............. UTIlSTY HOIST .......................... HOIST CONTROL PAN?Z ............

1-60

p~

-

-

p

-

~-

Hoist Power Switch and Indicstor Light ............. ...... , . Hoist Master Switch.. ... .. Hoist Control Switch ,.. ......,,, Cable Shear Switch . ... ...

. ........ .. . . .. ..... .

Power Switch

Li@

viii

1-60 1-62 1-62 1-62 1-62 1-62 1-63 '

.,, ..... ........ . ........ .....

1-63 1-63

,

1-63

,

,

Hoist Control Switch . Hoist High Temperatare Warning

rae-

1-60 1-60

,

.-......,.....-.....-...--1-63

........... .

s~vitche~.

Part 2 Table of Contents

FLIGHT MANUAL TABLE OF CONTENTS (Cont) SECTION

PAGE

LOADING INFORMATION .......................................................

2-1

MP W I HTAN LOCATION................................................................................

2-1

FOR AIRCRAFT WITHOUT THE UTILITY HOIST OPTION INSTALLED ........................................................................... FOR HELICOPTER OPERATION WITH THE UTILITY HOIST OPTION INSTALLED BUT NOT OPERATED DURING FLIGHT .................................................................

I OAnlNG INSTRUCTIONS...................................................... LOADING INFORMATION - UTILITY HOIST LOADS ............ C D 1 . CENTER OF GRAVITY LOCATION......................................... LOADING INSTRUCTIONS...................................................... NOISE ABATEMENT..................................--.........-.................. NOlSF ABATEMENT PROCFf3URFS..................................... General Procedures Recommended to Minimize Acoustical I m ~ a c.....................................-.............,............... t Ap~roachProcedure................ .................................. Take-Off Procedure ............................................................... Enroute Procedure............................. . ..................................

. . . .

SUPPLEMENTAL PERFORMANCE DATA .............................

NOVEMBER 21,1978 Revised November 4,2002

Part 2 List of Illustrations

FLIGHT MANUAL

LIST OF ILLUSTRATIONS FIGURE

TITLE Recommended Personnel Exit and Approach ................. Sliding enger Door Open Lights ............................... Three View Dimensional Diagram.................................... Pilot's Compartment (Typical) ......................................... Center Console (Typical)................................................ Engine Control Quadrant.................................................. Cyclic and Collective Stick Grips ...................................... Instrument -Typical VFR Configuration................ Instrument . Typical IFR Configuration ................. Fuel Flow and Fuel Totalizer Indicators ........................... Engine Starting System .................................................... Master Switch . Typical (Aircraft Serial Numbers 760331 and Prior) .......................................................... Engine Air Particle Separator System (EAPS)................. Auxiliary Circuit Breaker ......................................... Master Switch . Typical (Aircraft Serial Numbers . . ................... 760332 and Subsequent) ...................... Manual Rotor Brake Master Cylinder ............................... Fuel System ...................................................................... Fuel Crossfeed Operation ................................................ Engine Prime.................................................................... Fuel Quantity Data ........................................................... Fuel Gage Test and N i K s Digit Light Switch ........ Upper Circuit Breaker . DC Primary Buses Typical AC Generator System (Aircraft Serial Numbers 760270 and Subsequent) ................................................. Upper Circuit Breaker . DC Primary Buses Typical AC Generator System (Aircraft Serial Numbers 760001 thru 760269) ........................................................ Lower Circuit Breaker - DC Essential and AC Buses - Typical AC Generator System (Aircraft Serial Numbers 760270 and Subsequent) ................................. Lower Circuit Breaker - DC Essential and AC Buses -Typical AC Generator System (Aircraft Serial Numbers 760001 thru 760269) ........................................ DC Electrical System........................................................ AC and DC Generator Overspeed Test s ............... Deleted .............................................................................

NOVEMBER 21.1978 Revised November 4. 2002

PAGE

Pan 2 9~m~~m/dE List of lllusmtions FUGHT MANUAL

FIGURE NUMBER

TITLE

PAGE

AC Electrical System .Typical AC Generator System ....................:...................................................... AFCS Phase 11 Control and Circuit Breaker s ............................................................................ AFCS Phase 111Control and Circuit Breaker s ............................................................................ Landing Gear Conrrol .............................................. Caution- Advisory and Master Warning s ................... Cockpit Floodiighr and Utility Light ......................... Overhead and Fire Extinguisher Control s :....,.......... Flotation System ................... . . ... .... .......................... Heater and ECL' Control s ......................................... Cargo Hook System ........................................................... Utility Hoist ....................................................................... Center of Gravity Chart ..................................................... Weight and Center ofGraviry Envelope ............................. .AircraftDiagram ............................................................... Weighing Procedure .......................................................... Standard 14 Place Corf~gumtion........................................ Standard Cargo Configuration ............................................ Chan A Typical ................................................................ Chan B , ............................................................................

.

Chart E .Cockpit and Cabin Companment Weight and Moment Table Ctran .............................................. Chan E .Usabie Fuel Weight and Moment Table .............. Cban E .Engine Oil and Windshield Washer Fluid Weight and Moment Tables ......................................... Cfian E .Internal Cargo Weight and Moment Table ............................................................................. Chan E .External Cargo Weight and Moment

Sliding Door Weight and Moment Table ........................... Sample Loading Conditions ...-................................... ....... Horizontal Center of Gravity Chan .....................-............. Weight and Horizontal Center of G ravity Envelope ...................................................................... Lateral Center of Gravity Qlan ........................................ Permissible Hoist Load for Combinations of Gross Weight and Lateral Center of Gravity ...................

NOVEMBER 2 1. I978 Revised March 2.1 989

3-5

3-6 3-7 3-8

Pan2

&~mm/~dE FLIGHT MANUAL

~ i sof t Illustrations

LIST OF ILLUSTRATIONS (Cont) FlGURE NUMBER

PAGE

Aircraft Diagram ............................................................... Weighing Procedure .......................................................... Standard llace Configuration, Crew. enger, and Baggage Centroids .....................-...-------Standard Cargo Configuration Cargo Centroids ................ Chan A ......................-....-.-........-----....-.---.-......------..-.--Chart B ...................................... *.................................. -.-. Chan C .............................................................................. Chart E Cockpit and Cabin Compartment Weight and Moment Table ......................................................... Chart E - Usable Fuel Weight and Momen~Table (Gravity Fueling) ......................~~~-~.~....~.~.~.................-.. Chan E - Engine Oil Weight and Moment Table ................. Chart E - Windshield Washer Fluid Weight and ....----.............--Moment Table ..................................... Chan E - internal Cargo Weigh~and Moment Table .............................................................................. Chart E - External Cargo Weight and Moment Table ......................................................... ......----.-..-.. ---Chan E Utility Hoist Load Weight and Moment -..---Table ....................................................................... Chart E - Sliding Door Weight and Moment Table -.--..--..-. Chart E - Hoist Operator Weight and Moment Table ............................................................................. Sample Loading .....-..-.............-.-. ...--........-......------..

-

-

-

-....

xii

NOVEMBER 21; 1978 Revised March 2,1989

FLIGHT MANUAL

Part 2, Section I Description

The Model S-76A, manufactured by Sikorsky Aircraft Division of United Technologies Corporation, is a twin engine, single main rotor helicopter designed to cany up to 13 engers and a pilot. Flight controls and instrumentation for a second pilot are also available as optional equipment. Various cabin configurations are available and range from the standard interior arrangement to a four- to eight-place executive version. Four large doors provide easy entrance and exit from the cabin. Baggage is stored in a compartment aft of the cabin. The four-bladed main rotor is mounted on the main gear box, which is directly above the cabin. The two Allison Gas Turbine 250-C30 or 250-C30S engines are mounted side-by-side aft of the main gear box. Both engines have separate drive inputs to the main transmission, which reduce engine rpm and distribute torque upward to drive the main rotor, and aft through intermediate and tail gear boxes to drive the four-bladed tail rotor. Flight control forces from the main and tail rotor blades are reacted by a two-stage hydraulic servo system. The main rotor flight controls have a stick positioning and force gradient system. The tricycle-type landing gear is hydraulically retractable. Dimensions are shown on Figure 1-1.

Four doors, two on each side of the fuselage, are provided. Entrance to the pilot's compartment and the first row of enger seats, when equipped with the standard three rows of enger seats, is through the left and right forward doors. Entrance to the second and third rows of enger seats is through the left and right aft doors. GUIDELINES FOR ENGER BOARDING AND EXIT. CARGO LOADING AND GROUND OPERATIONS WITH ENGINES AND/OR ROTORS TURNING Turning rotors represent considerable hazard to any persons in their vicinity. The best method for prevention of rotor strike incidents are training of flight and ground crew and effective briefing of engers. Because well-trained and briefed persons can become preoccupied and helicopter operations may involve areas where access control is limited, special attention must be given to the particular capabilities and characteristics of the S-76 whenever the engines andlor rotors are turning. MAIN ROTOR CLEARANCE The minimum clearance of the main rotor at 107% Nr and flat pitch is nominally 8 feet at the front of the helicopter. This clearance, while normally sufficient for safe approach and departure can be reduced significantly by a number of

NOVEMBER 21,1978 Revised November 4,2002

Part 2, Section I Description

FLIGHT MANUAL

factors. These factors include control position, rotor speed, winds, and terrain variation.

ENTER

I

=wiA

Figure 1-0. Recommended Personnel Exit and Approach

NOVEMBER 21,1978 Revised November 4,2002

SA 4047-76-1

~///~RHK(~Y/TE 2, Part

FLIGHT MANUAL

Section I Description

TAIL ROTOR CLEARANCE The minimum clearance of the tail rotor is nominally 6 feet by 6 inches. Access to the tail rotor is somewhat limited by the horizontal stabilizer, but ground clearance may be reduced greatly by uneven terrain. ROTOR BRAKE The S-76 rotor brake is capable of holding the rotor with both engines running at ground idle. This allows convenient loading and unloading without the necessity of stopping engines. Persons must not be complacent about the possibility of a rotor strike since brake slippage or release will allow the rotor to turn. Treat the rotors as if they are turning whenever the engine(s) islare running. GUIDELINES AND CAUTIONS The following guidelines are useful in planning operations and are a mandatory minimum. Other appropriate guidelines should be considered when conducting crewlenger briefings and when writing operations manuals and instructions. Never approach or depart the helicopter unless the pilot sees. acknowledges your intention and approves. Approach only from the sides of the aircraft (see Figure 1-0). Never approach from the front or rear. WARNING Some persons may have prior experience with helicopters for which approach from the front is preferred. Approach from the front of an S-76 is not permitted. It is essential that all persons are properly informed regarding the safe approachldeparture angles of the S-76. Ensure the cyclic control is neutral and the trim is ON. Do not leave the controls unattended. Use ground guides whenever possible to assist engers. If guides are not available, use available flight crew. Never move under the rotor unless it is stopped or turning at flight speed. Always treat the rotor as if it is turning. If practical, position the aircraft so that the natural approach and departure lanes are from the sides of the helicopter. Carefully observe the landing area terrain. Ensure that engers and crew are not moving over rising or irregular terrain. NOVEMBER 21. I978 Revised ~overnbe;4,2002

1-2A

Part 2, Section I Description

I

9.

FLIGHT MANUAL

Standardized procedures will reduce the likelihood of persons being surprised by rotor or engine starts and stops.

HINGED DOORS

Outside door handles, at the aft end of each door are lifted to open the doors. Inside door handles are at the rear of the forward doors and at the front and rear of the aft doors, for convenience from either row of seats. The inside handles are pulled aft to open the doors. Safety lockpins in the top aft corner of the doors prevent accidental opening of the doors from inside the enger compartment. After the doors are closed and latched, the safety lockpins are engaged by pressing lockpin buttons below the window frame of each door. Engagement can be visually checked by observing that the bottom of the lockpin moves upward into the frame (no red showing). Before opening a door from the inside, the lockpin must be retracted by pulling up the lockpin button. When the door is opened from the outside, the lockpin is retracted when the door latch is lifted. Electric Door Locks

Electric door locks may be installed as optional equipment in the enger doors. The door locks are controlled by an ELECTRIC DOORLOCK switch on the center console, with positions marked LOCK and UNLOCK. When the. enger doors are shut properly, activation of the door lock switch will seat the door lockpins. The door lock buttons on the enger doors will still manually lock and unlock the enger doors. The electric door lock circuit is connected to the essential bus by a circuit breaker marked DOOR LOCKS.

s

i

1,

'!I

I!

I

CAUTlONlADVlSORY

Figure 1*A.

I INSTRUMENT

Sliding enger Door Open Lights

NOVEMBER 21,1978 Revised November 4,2002

FLIGHT MANUAL

Part 2, Section I Description

SLIDING DOORS (OPTIONAL) Sliding doors may be installed in place of the right or left hinged-type enger doors. After two primary and two secondary latch pins are retracted, the door opens by moving it out from the fuselage and then aft along the fuselage. The sliding door cannot be jettisoned, but is removable for maintenance. The door provides entrance and exit for engers in the center and aft rows of seats. The door has a large window that can be jettisoned from the inside. Operating controls are inside and outside handles, inside and outside locks, and an inside window jettison handle. Both handles are installed in recesses near the center of the door. To open the door from the inside or outside requires that the lock be turned to the UNLOCK position after which the handle is turned down. A window jettison handle is installed in a recess at the center of the door below the window; it is covered by a readily removable transparent cover so it will not be moved accidentally. When the cover is removed and the yellow knob pulled out and forward, the window can be pushed out of the door. The secondary latches in the top and bottom of the door are backups for the primary latches in the side of the door and prevent the door from being opened accidentally from the inside. The position of the lock in the locked position and the handle in the closed position provide a visual means for determining proper latching of the door. A microswitch on the lower secondary latch also indicates to the crew via the caution-advisory that the door is closed and locked. The door slides on an upper and lower track and swivel assembly. When the door is pushed open, the door and tracks move out from the fuselage about 4 inches. As the door slides aft, the track and swivel assemblies extend; when they reach full extension a stop within each assembly halts further extension and a spring-loaded hold-open stop on the door engages a catch to prevent the door moving forward. To close the door, the outside handle must be pulled slightly forward to disengage the hold-open stop. Flight operations may be conducted with the r.h. sliding door open on helicopters modified by kit PIN 76060-200015-011 within the limitations contained in Part 1, Section I.The sliding door is held in the full open position by a self-locking detent. To close the door, turn the door handle to release the door from the detent and slide it to the closed position and lock. With sliding door open, RHDR annunciator, on the instrument , will light.

Two Allison Gas Turbine 250-C30 or 250-C30S engines, rated to 650 shaft horsepower each, are mounted aft of the main gear box. The power turbine of each engine is connected to the main gear box with separate input shafts. The main gear box input shafts contain the freewheel units. The engine fuel control unit regulates power output by controlling gas producer speed. Gas producer speed (NI) is established by the power turbine governor which senses power turbine speed (N2) change in response to load (collective stick position). A collective bias system minimizes N2 (and rotor) speed changes through variable power requirements. The DCA (Droop Compensator Amplifier) minimizes rotor RPM droop during large power demands from low power setting (e.g. power recovery autorotation). The DCA is armed when N1 is below 75%. When a large power demand is made (increasing collective more than about 1 inch-perNOVEMBER 21,1978 Revised November 4.2002

Part 2, Section I Description

FLIGHT MANUAL

second), the DCA forces the collective bias meter of each engine to the upper limit and closes the heaterlECU bleed valves. These two actions result in quicker engine response with less transient rotor droop. Either of the two DCAs will cause both engine collective bias motors and bleed valves to respond. The collective bias motors and bleed valves will reset upon completion of the transient condition. The collective bias systems for the No. 1 and No. 2 engines are connected to the No. 1 dc primary and No. 2 dc primary buses, respectively, by circuit breakers marked CLTV BIAS. The DCAs are powered through the CLTV BIAS circuit breakers. Ice protection is provided by a bleed-air anti-icing system. An air particle separator to protect against FOD is optional. The engines are started electrically using starter-generators.

NOTE When the helicopter has been operated in a corrosive atmosphere, such as salt water spray, ensure that an engine wash is performed. ENGINE RATINGS AND RECOMMENDED USAGE The model 250-C30 and 250-C30S engines are FAA certified to the limits defined in Figure 1-5 of Part 1. However, in order to conform to the engine manufacturer's warranty policy and to prolong engine life, power settings should be limited to 694°C T5 (Normal Cruise Power) in lieu of maximum continuous power except as required to accomplish the following operations and checks: takeoff, landing, and hover (power levels up to 768°C T5 for up to 5 minutes) actual OEl (not to exceed 826°C Ts, for 2% minutes; and 798°C T5 for 30 minutes) power assurance checks as defined in SECTION II, NORMAL PROCEDURES fuel control limiting check as described in the S-76 Flight Check Procedures Manual conditions where power required to prevent aircraft damage or personal injury exceeds that provided by normal cruise power The highest values of torque and T5 conforming to the engine manufacturer's warranty policy for continuous operation are the normal cruise values shown below. The region between normal cruise power and maximum continuous power is defined by the green and white arcs on the torquemeter and T5 indicator and by the green and white light on the digital Ts repeater. ENGINE FUEL SYSTEM Fuel from the aitframe fuel system is initially down into the low-pressure chamber of the jet inducer by the high-pressure gear stage of the fuel pump. By return from the high-pressure stage of the fuel pump to the jet inducer nozzle NOVEMBER 21,1978 Revised November 4,2002

SA 4047-76-1

A~R~~B~BEM~~K Part 2, Section Description

I

FLIGHT MANUAL

induces secondary flow and increases fuel pressure to 15 to 25 psig. This low pressure fuel from the jet inducer then es through a filter and enters the high-pressure gear stage of the fuel pump. Fuel under fuel pump pressure is then delivered to the gas producer fuel control. The gas producer fuel control meters fuel to the combustion section as required, to produce a constant power turbine speed at the power required by the rotor and drive system. The power turbine speed selected with the N2 trim switches is automatically maintained by the power turbine governor's action on metered fuel flow to the gas producer. The overspeed protection system prevents N2 from exceeding safe limits by reducing fuel flow the gas producer to a minimum, if an overspeed should occur. ACCELERATION BLEED VALVE A bleed valve on the front of the engine compressor scroll is programmed to unload the compressor during certain conditions to protect against engine acceleration surges or stalls. The valve is scheduled through compressor discharge air pressure and electrically sensed engine inlet air temperature. At inlet temperature above 15"C, the valve is closed for all engine operating conditions. Below 15"C, the valve is open up to about 83% NI and closes at higher power. Because the valve is closed by compressor discharge pressure, the exact NI at valve closing will depend on the operating altitude, with higher closing NI at higher altitudes. Because of differences in bleed valve schedules between engines, a mismatch of engine T5 might be noted under certain cruise conditions, especially at 12" to 15°C OAT. ENGINE OIL SYSTEM Each engine has an independent dry sump oil system with separate oil coolers and tanks between the engines. Oil capacity is about 1.25 US. gallons (1.04 Imperial gallons or 4.73 liters) per tank. Oil flows from the tank to the enginedriven oil pump, which distributes oil under pressure to lubricate the various components of the engine. An oil filter, filter by, and a pressure regulating valve are incorporated into the system. Magnetic chip detectors, installed at the bottom of the engine accessory gear box and at the engine oil outlet, light a caution light when metallic chips are detected. A dual gage on the instrument indicates oil temperature and pressure, and a caution light indicates low pressure. An engine oil pressure torque sensor is used as a means of indicating engine output torque.

NOVEMBER 21,1978 Revised November 4.2002

Part 2 Section I -

Description

FtKiCCT MANUAL

I

Figure 1-1. TirlPe-View Dimensional I)iagam

NOVEMBER 21, 1978 Revised JUNE 28, 1984

1-3

NCHT MANUAL

-

AFCS INDICATOR



AFCS NO. 2

KTROL /W

/ PARJUNG BRAKE HANDLE

I

Figure 1-3. Carter Console (T9pSa)

NOVEMBER 2 5 1978 Revised July 26, 1982

.

1-5.

E N G m CONTROLS EngiPe Levers

No. 1and No. 2 engine levers on the engine control qrr;rdrant (Figare 1-41 are connected to the fgel control by mechanical linkage. The quadrant has three marked positions, OFF, IDLE, and FLY. Detents at each positian indicate proper positioning of the lever. The enpine lever schedules fuel flow and gas producer speed in the OFF to GFtD IDLE range, and establishes the N1 speed limit for W m n engine power in the FLY position. In addition to the mannal positioning of the lever, it will also be moved aft to OFF when the respective fire-her T-handle is pulled. A starter button is in the grip of each lever. N2 Speed Trim Switches No. 1and Xo. 2 engine N2 speed trim swEtches on the grip of the collective pitch lever (Figure 1-5) m used to establish ihe desired power turbine speed and t o match engine torques. is a&omatically maintained by the power turbine governor's action on metered fPel to the gas producer. The switches are marked ENG TEUM 1and 2. To increase N2, the switch is moved to + (fwd); to deczease N2, the switch is moved to (aft). Trim range (full aft to full fomard) is about 96% N2 to 107%N2. If copilot's controls are izrstalled, the pilot's switch w i l l override any input from the capilot's switch. A collective bias actuator and a collective pitch signal zesponds to collective slick movement and resets the N2 governor to mainlab a canstantrotorrpmasestablishedbythe~speed.trimwllencollective pitch is increased o r decreased. The No. 1and No. 2 engine N2 speed trim

-

-

NOVEMBER 21, 1978 Revised JCME 28, 1984

Part 2 Section I Desc*tioo RIGHT MANUAL

system is connected to the dc esseatial bus by cireuit breakers marked SPD lTUM, and the collective bias system is corrnected to the dc essential bas by a cizroit breaker marked CLTV BIAS.

Toquemeter A dual torquemeter marked PER* TORQUE has pointes marked 1 and 2, which indicate No, 1 and KO. 2 engine t o m OUQUL. The torquemeter sensing system, within the engine accessory gear provides a hykwdic signal which is directly proportional to torqtie outpt. This signaf is tansmi.tted electrically to the torquemeter on the instrrrment . The electrical circuits for *.No. 1 and KO. 2 engine torquemeters are connected to the No. 1 and No. 2 dc primary buses respectively, by circuit breakers

marked TORQ.

NOVEMBER 21, 1978 Revises July 26, 1982

-2

Section1

Description

w CYCUC mtK GRIP

COLLECllVT STICK GRIP

Frgr~e1-5,

clic and Collective Sick Grips

NWEMBER 21, 1978 Revised M y 26, I982

NO21, 1978 Rwised FEBRUARY21, 1979

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 1 1. 12. 1 3. 14. 15. 16. 17. 17A. 10. 19. 20. 21. 22. 23.

LANDING GEAR UP WARNING LIGHT AIRSPEED INDICATOR DUAL TORQUEMETER DIGITAI. 15 REPEATER TRIPLE TACHOMETER Vne POWER-OFF PLACARD ATTITUDE INDICATOR DIRECTIONAL GYRO INDICATOR TURN-AND-SLIPINDICATOR AUDIO CONTROL MASTER WARNING ALTIMETER VERTICAL VELOCITY INDICATOR RADAR ALTIMETER INDICATOR CLOCK FLOATS TEST (OPTIONAL) NO. 2 VHF COMMUNICATION CONTROL (OPTIONAL) FUEL TOTALIZER TEST - DC GENERATORS CAUTION/ADVISORY Vne POWER-ON PLACARD NO. 1 VHF COMMUNICATION CONTROL (OPTIONAL) NO. 1 VHF NAVIGATION CONTROL (OPTIONAL) N1 TACHOMETERS

21. 25. 26. 27. 28. 29. 30. 3 1. 32. 33. 34. 35. 36. 37. 30. 39. 40. 41. 42. 43. 44. 45. 46.

15 INDICATORS

FUEL PRESSURE INDICATORS OR FUEL FLOW INDICATORS ENGINE OIL PRESSUREAEMPERATURE INDICATORS DUAL HYDRAULIC OIL PRESSURE INDICATOR MAIN GEAR BOX OIL PRESSUREAEMPERATURE INDICATORS DUAL FUEL QUANTITY INDICATOR FUEL GAGE TEST AND N1A5 DIGIT LIGHT SWITCH LANDING GEAR CONTROL CLOCK MASTER WARNING LANDING GEAR UP CAUTION LIGHT AIRSPEED INDICATOR DUAL TORQUEMETER DIGITAL 15 REPEATER TRIPLE TACHOMETER AUDIO CONTROL ATTITUDE INDICATOR DIRECTIONAL GYRO INDICATOR TURN-AND-SLIP INDICATOR ALTIMETER VERTICAL VELOCITY INDICATOR RADAR ALTIMETER (OPTIONAL) Vne POWER-OFF PLACARD

NOTE ALL COPILOT'S INSTRUMENTS ARE OPTIONAL

V,

9

E

2

?'

Part 2 Section I Description

-.. ... ..- "

-

FUGHT MANUAL

--

Fi,aure

-

-

1-7. Instnrment Typical IFR Confi,~uration (Sheet 1 of 2)

NOVEMBER 21, 1978 Revised JANUARY 17, 1983

1. 2. 3. 4. 5.

6. 7. 8. 9. 10. 1 1. 12. 13. 14. 15. 16. 17. 18. 19. 20. 20A. 21. 22. 23. 24. 25. 26. 27. 28. 29.

LANDING GEAR UP WARNING LIGHT AIRSPEED INDICATOR DUAL TORQUEMETER DIGITAL 15 REPEATER TRIPLE TACHOMETER Vno POWER-OFF PLACARD VOR MASTER SLAVE ATTITUDE INDICATOR RADIO MAGNETIC INDICATOR (OPTIONAL) COURSE DEVIATION INDICATOR (OPTIONAL) AUDIO CONTROL MASTER WARNING ALTIMETER VERTICAL VELOCITY INDICATOR RADAR ALTIMETER INDICATOR CLOCK FLOATS TEST (OPTIONAL) MARKER BEACON (OPTIONAL) NO. 2 VIiF COMMUNICATION CONTROL (OPTIONAL) VHF NAVIGATION CONTROL (OPTIONAL] FUEL TOTALIZER TEST GENERATORS CAUTION/ADVISORY ADF CONTROL (OPTIONAL) WEATHER RADAR DISPLAY (PROVISIONS) Vno POWER-ON PLACARD DME INDICATOR CONTROL (OPTIONAL) ATTITUDE INDICATOR NO. 1 VHF COMMUNICATION CONTROL (OPTIONAL) NO. 1 VHF NAVIGATION CONTROL (OPTIONAL)

-

30. 31. 32. 33. 34. 35. 36. 37. 38. 39, 40. 4 1. 42. 43. 44. 45. 46. 47. 49. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58.

ATC TRANSPONDER CONTROL N i TACHOMETER 15 INDICATORS FUEL PRESSURE INDICATORS OR FUEL FLOW INDICATORS ENGINE OIL PRESSUREflEMPERATURE INDICATORS DUAL HYDRAULIC OIL PRESSURE INDICATOR MAIN GEAR BOX OIL PRESSURE/lEMPERATURE INDICATOR DUAL FUEL QUANTITY INDICATOR FUEL GAGE TEST AND NI/PTIT DIGIT LIGHT SWITCH LANDING GEAR CONTROL CLOCK VOR TEST (OPTIONAL) DME SELECTOR MASTER WARNING MARKER BEACON (OPTIONAL) AIRSPEED INDICATOR DUAL TORQUEMETER TRIPLE TACHOMETER AUDIO CONTROL VOR MASTER/SLAVE (OPTIONAL) ATTITUDE INDICATOR (OPTIONAL) RADIO MAGNETIC INDICATOR (OPTIONAL) COURSE DEVIATION INDICATOR (OPTIONAL) LANDING GEAR UP CAUTION LIGHT ALTIMETER VERTICAL VELOCIN INDICATOR DIGITAL T5 REPEATER RADAR ALTIMETER (OPTIONAL) Vno POWER-OFF PLACARD

NOTE ALL COPILOT'S INSTRUMENTS ARE OPTIONAL

Part 2 Section I DescriDtion ----~Z

FLIGHT MANUAL

Ts (Power Turbine Inlet Temperature) Indicators I TWOdial pointer T5 indicators marked Tg°C X 100 a r e on the instrument

1 . Thermocouples in the gas producer turbine of each engine sense

temperature changes and electrically transmit them to the indicators. Each indicator has a redundant electrical power supply. The So. 1 engine T5 indicator i s powered by the No. 1 dc primary bus and the dc essential bus through two circuit breakers marked Tg. The No. 2 engine Tg indicator is powered by the No. 2 dc primary bus and the dc essential bus through two similar circuit breakers. Digital T s Repeater A dual digital T5 repeater marked T5, 1-ENGINE-2, "C is on the instrument . The repeater consists of three-digit readouts and five colored indicator lights for each exine. The digital readout displays the same value as its correspondanalog Tg indicator. The colored lights go on to advise of the Tg operatikff temperature range. The relationship of the lights to T5 operat- temperatures is shown on the Instrument m e Markings Illustration in Part 1. Intensity of the digital readout and lights the knob marked N1/T5 DIGITS on the fuel gage may be adjusted by test (Figure 1-14). The indicator lights a r e tested by pressing the caution- advisory TEST button. An additional digital T5 re-peater is installed on the copilot's side.

N1 (Gas Producer) Tachometers

-

-

Two N1 ,oas producer tachometers marked TACH % X 10 are combination dial-pointer and digital readout -type indicators. The digital readout indicates N1 as a two digit number and a decimal (99.5). I£N1 should go over 100%the values will appear as 00.0 for loo%, and 01.5 for 101.5% etc. Light intensity of the digital readout may be adjusted by the knob marked Nl/Tg DIGITS on the instrument . Magnetic pickups sense changes in W 1 speed and transmit these changes electrically to the indicators. Each en,$ne tachometer has a redundant power supply. The No. 1e k g e Nl indicator is connected to No. 1dc primary bus and the dc essential bus by two circuit breakers marked S1. The No. 2 engine N1 indicator is connected to the KO. 2 dc primary bus and the essential bus by the circuit breakers marked N1. Triple Tachometer A triple tachometer marked PERCENT R% indicates the N2 (power turbine speed) of each en,gi.ne and Nr (rotor speed). The outer scale marked ENG and the 1and 2 pointers indicate the K2 of the No. 1and No. 2 engges. The inner scale marked ROTOR and the R pointer indicate N,. Magnetic in K2 or Nr speed and transmit these c w f f e s pickups sense -es electrically to the indicator. The electrical circuits for No. 1and No. 2 engine N2 are connected to the No. 1 and No. 2 dc primary buses by circuit breakers marked N2. The Ei,tachometer has a redundant power supply, be-u connected to the No. 2 dc primary bus and the dc essential bus by circuit breakers marked NR.

NOVEMBER 21, 1978 Revised JONE 28, 1984

F U M MANUAL

Part2SectimI Description

Fuel Pressure Indicators

Two fuel pressure Micators marked FUEL PRESS-PSIX 10, indicate fuel pressare at the fuel filter assembly. The No. 1 engine fael pressare indicator is connected to the No. 1 dc primarp brrs by a circait breaker marked FUEL PRESS. The No. 2 engiw fuel pressare indicator is cannected to the No. 2 dc primary bus by a circuit breaker marked FUEL PRESS.

On some heiicopters, fi& flow indicators and a fuel tatalizer are installed on the instramat in lieu of foel pressnre gages, This system indiCafgb the rate at which fuel is being consumed and the total feel.zemaining inthefueltanlrs. Flowmetersinthemainfuellinetoeacheaginec~ turbine wheels which rotate at speeds proportional to the pulse signals that are directly proportional to the volume of fuel flow. These pulses are cmditioned by a computer auit into a signal whi& drives the fuel flow indicators (one per engine) on the instrmnent . The dial of each f u e l . h .indicator is marked FUEL FLOW, LBS/HR. The pulses are also combined by the computer unit to drive a tatalizing indicator ahich reflects the fuel re-inthefuelsystem, Thgdialofthefueltoblizerism&&FUEL REbdAIMNG POUNDS. The tatalizer is inaccmate if one or krth flowmeters are inopeative. Four preset switches an the face of the f& totalizerindicatorareosedtosetthetatalf&qoantitycontainediathe fuel tasks as indicatedon the dnal fwd qrrarttfty gage. Eachpresetswitch is pressedinasaEtenaszeqPired~~digTtsanthe~t4talizerinditate

-

totalipel, EachclicBofthepresetswitchsmovestheapplicaUemrmber~

one digit.. The No. 1 engine fuel flow is connected to the No. 1 dc p r i m a r p b u s b p a c i r c r r i t b x w k e r ~ d N O . l E N G F W 3 L F L O W , andthe No. 2 enginefuel5owsyskmis connectedtotheNo. 2 deprharybusby acirtrdtbreakermadcedNO.2ENGFUELFIX)W. Theseci~~&breagers aze hskdled in lieu of the FUEL PaESS cirertit breakers.

NOlEMBEX 21, 1978 Revised JLTNE 28, 1984

a r t 2 Section I Descriufion

The 61 ENG FUEL PRESS or #2 E N 'FUEL PFESS warning light an the caatian-advisory ( F i 1-25) goes on when fuel pressure drops below 5 psi. Each low fpel pressure circuit is connected to the dc essential bus by separate circuit breakers marked FUEL PRESS WARh'.

Engine O i l Pressare/Temperatur&Indicators Two dual indicators indicate engine oil pressure, as regdated by the oil pump and pressare revalve, and engine oil inlet temperature. The dial of each i s marked ENG OIL, X 10. The left scale is marked T (temperatPre)a t h e rightscale i s marked P (pressure). The Bio. 1 engine pressure t e m p e m indicator is connected to the No. 1dc primary bus by circuit breakers marked OIL-TEMP and PRESS. The Bio. 2 engine pressmet e m p e m indicator is connected to the No. 2 dc primary bus by the circuit breakers marked OIL-?TEMP and PRESS

Engine Oil Low Pressme Warning Light The #I ENG OIL PRESS or #2 ENG OIL PRESS warning lights on the cantionadvisory (Figure 1-25) go on when engine oil pressure drops below aPPr&maWy 48 psi. Each low pressate circuit i s connected to the dc essenlial bas by separate circuit breakers marked OIL PRESS WARN.

The %1ENG CHIP or #2 ENG CHIP caution lights on the caution-advisarp (Figure1-25) go on whea metal chips =detect& in the lower part of the engine accessory gear box or the engine oil mU& After a light is switched on by the detector unit, an electrical holding cirtait willkeep the li@t on until electrical power is removed by pPlling the apprapriafe circuit breaker, or rmtil the eLectrical system is shut down. Each chip detector circaitisco~nectedtothedc~busbyseparartecircuitbrealrets marked CHIP DET.

NOVEMBER 21. 1978 'RevisedFEBRUARY 22, 1985

P;ut2SectionI Description

Engine Bleed-&

Advisory Light

The EM; BIZED AIR advisory Light on the cautim-acivisory (Figme 1-25] will go on when any of the systems wing compressor discharge air ate operating. Systems usiug compressor discharge air are; the engine air particle sepaator system, cabin heater, and eavironmental control systgm.

ENGZNE STARTER AND I G r n O N m

M

(Figpre 1-8)

Starter/geneators provide torqae for engine starting when supplied with a dc power soarce. After starting, the starter/generators are driven by the m e aad are the primary source of electrical power. The starter/ generators, mounted on the engine accessory gear box, are connected to the dc essential bus when they function as starters, Electrical p e r for engine starting ma? be provided by the battery, external power or from the opposite stazter/generator, once the first engine is started. Controls for cansist of a master start switch and a starter switch for each engine. Starter control cirmits are connected to the dc essential bus by Ume circait breakers marked SIB CONTROL, on the lower circait breaker

-

m.

Master Start Switch T h e ~ S T A R T s w i t c h a n ~ ~ s w i t c h ~ ( F i 1-9)hastwo g P r e positions marked ON and OFF. ON provides dc sbrkr p o m r to ihe i d ridual&cuB which are then controlled by the engine starter buttas. OFF removes power from starter circuits,

The starter butttms in the b o b of eaeh N1 lever (Figure 1 4 ) prwide dc powertothestarterandjgnithsystemsforeachengine. T o s t a r t a n engine, themasterstartswitchmPstbeON, theengineleverOFF, andthe starter button depressed Aftez the starter is engai@, a holding cirit keeps the starter engaged for a s long as the starter button is depresseqwbn the engixte lever is h c e d from OFF. When the desired N1 is reached for a start, the engine lever is moved to IDLE. As the engine lever is actoanced, power is applied a-dy to the -an spstem to effect tke start. When the batton is released, both the starter and ignition circuits will be &-energized, Ignitim is provided only daring -st=-!2cfe-

ofthe idlet Hot compressor discharge air is used for --ice guidevanesandfmntbearing~. &-=tube, colme-d to the compressor section, carrigs hot air forward and into the hollow campressor inlet guide vanes. The spstgm is controlled by a solenoid-

Part 2 Section I Description

, DC ESSENTIAL BUS

I

-

NO. 1STARTER BUlTON

NO. ISTARTER CONTROL RELAY

1 IGNITION^ NO. 1 EXCITERS

-

, -

NO. 1ENGINE START OR

L

DC INPUT FROM BATTERY. EXTERNAL POWER OR DPWSTTE GENERATOR

MASTER

NO. 2 STARTER BUTTON

NO. 1STARTER/ GENERATOR

8

I

MASTER START OR 11 w+H MGiiR SWITCH ON. MASTER START OR COMPLETESA CIRCUIT BETWEEN THE DC ESSENTIAL BUS (STARTER POWER SOURCE) AND ENGINE START ORS. NO. 1 AND NO. 2 ENGINE START ORS 1. WHEN ACTUATED BY THE STARTER CONTROL RELAYS. ENGINE START ORS CONNECT A CIRCUIT BETWEEN MASTER START OR AND ENGINE STARTERS. NO. 1AND NO. 2 STARTER CONTROL RELAYS 1. WITH ENGINE L N E R OFF AND STARTER BUTlON DEPRESSED, THE STARTER CONTROL RELAY CLOSES. THE CIRCUITTHROUGH THE ENGINE START ORS. CONNECTING THE MASTER START OR TO THE ENGINE STARTERS. 2. PROVIDE A HOLDING CIRCUIT AFTER INITIAL STARTER ENGAGEMENT WHICH PERMITS THE ENGINE L,WER TO BE MOVED TO IDLE WITHOUT STARTER DROP-OUT. 3. COMPLETE A CIRCUIT FROM DC ESSENTlAL BUS TO IGNITION SYSTEMS WHEN STARTER BUTTON IS DEPRESSED AND ENGINE L N E R IS MOVED T O IDLE.

Figure 1-8. Engine Starting System

NOVEMBER 21, 1978

1-15

Parf 2 Section I

-

Description

FUGHT MANUAL

TEST PiTOT HEAT CO -PILOT PILOT

EXT POWER

E R@ ON

-

sRcH,

OFF

EAPS ON

ON

ON

STOW

J

RESET

I

;6~6 HEATER CO.PILOT PILOT

RESET

r

MASTER START

@ i:iE! ON

RE L

ARMING

6 y5 116 ; ;69 ANTI COLL

POSITION

WASHER

WIPER

SLOW

I

OFF

OFF

OFF

n s n - s s ~ a g( c c ~ )

Figure 1-9. Master Switch , Typical

NOVEMBER 21,1978 Revised September 7,1990

-

-

operated valve in this line which controls an &-ice air valve. The sole- - anti-icing if an electrical maffimcnoid valve is mrmaLfy apen, tion occurs, and must be energized electricaLly to close. Tg will increase about 10"C when the anti-ice system is in use. The system for No. 1engine operates from the No. 1dc .iprimary bas and is protected by a breaker marked AI CONT. The system for No. 2 engine operates from the No. 2 dc primary bas and is protectedby a circuit breaker marked A1 mhT. Engine &-Ice

I

Switches

?tPo switches on the master switch (Figure 1-9) marked ENG ANTIICE No. 1 No. 2 have marked positions ON and OFF. Each switch cantrols the solenoid valve in the engine anti-icing air tube of the corresponding engine. ON provides anti-icing by opening the vafve (deenergizing the solenoid). OFF closes the valve (energizing the solenoid).

-

-

The fi and and2 ENG ANTI-ICE caution hgks on the ~ - a d v i s o r g (Figure 1-25) w i l l go on when an engine anti-ice pressure sv&h does not sense air pressure (valve does not fully upen) when the corresponding engine &-ice m h is ON. The cadion light and the ENG ANIII-ICE ON advisory light will also go an when an engine anti-ice pressare srPitch senses air .pressure (valve does not fully close) when the corresponding anti-ice switch is OFF. Anfi-Ice On

Lights

The el and #2 ENGANTI-ICE ON a d m r y light on the carrtion advisay (Figure 1-25) will go on whenever the torrespondiag engiue &-ice system is aperating regardless of the switch posdion. For normal operrr.tian (engine operating)the light will go on when the switch is placed to ON. If the anti-ice air valve does nat close when the s r i k h is placed OFF (enmoperating),t h e c o r r e ~ ~ a n d c a ~ l i g h t s w i l l b e a n ,

NOTE Some helicopters may bave both thg engjne anti-ice systems anci EAPS installed. *these helicopters EAPS may be used as described in the following paragraphs when the --ice systems are off. When the engine &-ice systems arein use, EBPS is arrtb maiuxlly Switched cng however, the EAPS swftch, the EAPS c;mtion and the EAPS advisory lights

domtamctiaa T h e a n t i - i e e ~ w l i g h b d t h e anti-iceacb5mxy~wXlimiicaeproper~pea-

~OfbaththeEBPSancianti-icesyskms. ENGINE AIR PARTICLE SEPARATOR (EAPS)

NOVEMBER 21, 1978 Revised JUNE 28, l984

Part2 Section1 Description

FLIGHT MANUAL

foengine c o m p r b e n t and is totally enclosed by the engine inlet cowl. The EAPS is acylindricalassembly incorporating miniatwe vortex'tubes arraaged radially between an inner aud oater wall .that form the scavenge chrcting. At the forward end, the EAPS has an inlet and i by doors that are opened and dosed by an eagine bleed-air operated actaator. When in the filter mode (EAPS-on) the duskladen air enters the vortex tubeswberea s w i r l is imparted to the air. The solid particles a r e centriluged to the outer wall and into the scavenge ducting together with a qoantity of ai.r equal to 8% of the engine air flaw. The scavenge blower is mounted in the Wet cowl together with the overbaard durting. The blower, pawered by the ac generator, induces the scavenge flaw and duxups it overboard along with the solid particles. The clean air es through an exit diff aud fnto the engine.

EAPS may be operated in a filter mode (ON) or in a by mode (OFF) as controlled by a switch on the master switch - One amber caution light and one green advisorg light on the caution advisory are incorporated into the system. The EAPS caution light warns the pilot of EAPS blockage, scavenge blawer failure, or tbat a discrepancy exists between the position of tbe EAPS by doors and the switch position selected. Five-second time delays are incorporated in the EAPS caution Ugbt to prevent nuisance lights during normal system a ~ o n s .The EAPS advisory light advises tbe pilot that EAPS is ope-. in the filter mode (ON). If an engine faits while EAPS is ope-, the system will autamatically dhengage EAPS on the remaining engine (open the by doors). If engine anti-ice systems are operating and an engine failme occurs bath the engine *-ice system and EAPS will cantisee to operate for the re-:engine.

-

One switch on the switch (Figure 1-9) marked EAPS coxxtrols theopemtimofEAPSforboth~s,TheswitchfrasmarBedpo~ ON and OFF. ON closes the EAPS by doors in each Wet air &ct and energjzes the e-y l operated blowers. OFF opens the by doors and shuts off the blowers. The sgstem for each engine is controlled by dc power fram tke No. 1and No. 2 dc primary buses recircuit breakers marked EAPS, The scavenge blowers operated on 3-phase a c power fram the ac monitor bas.

TheEAPS~ligfrtonthe~advia>rg~willligjltwhenthe E A P S s w i t c h i s p h c e d t o O N i f ~ b l g w e r f a i l s t o o p e z a t e iftheby , doors of either EAPS do not close or if EBPS blockage in the system causes excessive inlet air pressure loss. The Carrtim Wzt will aLso light ifthe by doors of either EAPS do not open when the EAPS switch is placed t o off.

NOVEMBER 21, 1918 Revised ATT27 7 0 7 0

Part 2 Section I Description FLIGHT MANUAL

EAPS Advisory Light The EAPS advisory light on the instrument will go on when the EAPS switch is ON and the system is functioning normally. If a malfunction is indicated by the EAPS caution light, the EAPS advisory light will go off. SNOW PROTECTICN KIT The sqow protection kit provides for continuous en,gine Qnition and protection for the engine while flying in a falling and blowing snow environment. Each engine inlet is provided with electrically powered heatelements on the firewall and engine bellmouth. These elements a r e normally powered from the primary dc power system when both dc generators a r e on the line. If either dc generator fails, the system is automatically powered from the ac generator. The auxiliary circuit breaker (Figure 1-9A(1)) is located on the pilot's side of the lower console. The snow protection kit operates in conjunction with the engine anti-ice bleed-air system, utilizing the same anti-ice switches on the master switch and the engine anti-ice caution Lights on the caution-advisory . Placing the engine anti-ice switches in the on position will provide bleed-air for engine anti-icing and electrical Dower for heating elements on the firewall and bellmouth. Jf =1and 82 ENG A~TI-ICE caution Lights do not go on, there is either a s n m protection system malfunction o r the ambient air temperature is too warm (about -lj°C OAT) to permit a valid check of the snow protection system. A control unit monitors sensors in the heatelements to automatically control their temperature. A failure of an inlet heater system is indicated by the respective engine anti-ice caution light. The snow protection kit also includes a snow shield and continuous ignition svitch. The snow shield .is a Teflon-coated aluminum sheet formed to fit around the engine inlet and attached to the forward transverse firewall. The Teflon layer reduces the surface adhesion when impacted by snow particles. The continuous ignition switch is utilized to provide a continuous engine ignition capabilie-. This switch allows the pilot to select continuous

-

"000 ENGINE Al-

/\

iNNO.

\/

1

'-

NO. 2

Figure 1-9A!1). Auxiliary Circuit Breaker NOVEMBER 21, 1978 Revised July 26, 1982

I 1

Part 2, Section I Description

FLIGHT MANUAL

ANTI-ICE SWrrCWLIGHT LOGIC TABLE (SNOW PROTECTION KIT INSTALLED) --

AWI-ICE CAUTION LIGHT (YELLOW)

SWlTCH POSITION

ANTI-ICE ON ADVISORY LIGHT (GREEN)

-

CONDITION Bleed-air pressure sensed in anti-ice system

ON

1

Anti-ice and snow protection kit systems operating normally

OFF

OFF

No bleed-air in the anti-ice system or simultaneous failure of anti-ice bleed-air and snow protection kit inlet heater System just turned on and not up to temperature or failure of snow protection kit inlet heater

ionition for takeoff, climb, and landing in falling and blowing snow conditions. The ;witch function is interlocked through the engine quadrant switches such that continuous ignition is disabled when the engine lever is OFF. ENGINE OVERSPEED PROTECTION SYSTEM

I

The engines are protected from overspeed by a mechanical overspeed system. If N? increases to about 118% the power turbine sovernor causes a reduction in fuel flow and engine power, to prevent overspeed.

NOVEMBER 2 1, 1978 Revised May 15, 1992

-

FUGHT MANUAL

Figure 1-9A

NOVEMBER 21, 1978 Revised AUGUST 27, 1979

Part 2 Section I Description

Part 2. Section I Description

FLIGHT MANUAL

TRANSMISSION SYSTEM

I

The transmission system consists of three gear boxes and associated shafting whi~.. transmit engine torque to the main gear box. Each engine is connected to the main sear box by freewheel units which permit single-engine operation or autorotation without drag from the inoperative engines. The main rotor drive shaft, to which the main rotor system is attached, extends upward from the main gear box. A drive shaft extends aft from the main gear box to the intermediate gear box at the base of the tail rotor pylon. Shafting extends up the pylon to the tail rotor gear box which drives the tail rotor. An accessory drive section at the rear of the main gear box drives the first and second stage hydraulic pumps, two main gear box oil pumps, the main gear box oil cooler blower, and the ac generatorTRANSMISSION OIL SYSTEM Pressure for lubrication of the main gear box is provided by two oil pumps mounted on and driven by the gear box accessory drive section. Oil is pumped from the sump of the gear box through an oil filter and then to the oil cooler. A thennostaticallycontrolled by valve directs the flow of oil either through the core of the oil cooler if the oil is hot, or through the by line if the oil is cold. The oil cooler blower is mounted on top of the gear box. The oil filler and oil level sight window are on the right side of the gear box. The intermediate and tail gear boxes are splash-lubricated and have no pressure or temperature g%es. !Main Gear Box Oil Pressureflemperature Gage A dual indicator marked XMSN OIL X 10 indicates oil pressure at the main gear box

oil inlet port and oil temperature in the gear box sump. The left scale is marked T (temperature) and the right scale is marked P (pressure). The temperature indicator is connected to the No. 2 dc primary bus by a circuit breaker marked XMSN TEMP. The pressure indicator is connected to the No. 1 dc primary bus by a circuit breaker marked OIL PRESS-XMSN. Main Gear Box Oil Low Pressure Warning Light

,

The MAIN XMSN OIL PRESS warning light, on the cautionladvisory (Figure 1-25), lights when oil pressure in the main gear box drops below about 14 psi. The system is connected to the dc essential bus by a circuit breaker marked XMSN OIL WARN-PRESS.

NOVEMBER 21, 1978 Revised May 15, 1992

F'art2 Section1 Description

RlCHf MANUAL Main Gear

Bmr Oil Hot Warniug Light

The MAIN XMSN OIL HOT warning light, on the caution adrisorg , goes over about 105*C. The ( F i i 1-25), lights when oil inlet system is connected to the dc essenkial bus by a circuit breaker marked XMSNoILwm-mp.

-

GEAR BOX CHIP DETECTOR/OIL HOT SPSTEM The MAIN CHIP, I h T GB CHW/aOT, or TAIL GB CHIPDOT cautinn ligfrts on the caatim-advisory go on when metal chips are detected in the oil sump of the respective gear baz t'FUZZ" burr16ffunits in each chip detector prevent k l s e indications of chips by burning off minute metallic particles (fuzz) tbai may collect an the detector anits and canse the caution lightstogoon. Theintermediateandkilgearbax~onlightSwill~o go on whenever oil t e m p e n goes over about 140°C. Once a light is switched on by the detector unit, a holding circuit w i l l keep the light on until electrical power is removed by pnlling the appropriate circuit breaker, or the electrical system is shut down. The chip detector/overtemp systems are connected to the dc essential bus by circuit breakers marked CHlP DETECTOR-MGB and IGB n;B.

-

The ROTOR BRAgE e t c h on the master saritch has positions marked ON-OFF-REL. The rator brage a not go on unless bath e levers are at IDLE or bebw and the N1 af at least one engine is at 59 or below (one ENG OUT warning light on). For normal the rotor bnLe should m t be applied unf5l romr speed decreu65%Nr. When the switch is placed ON under the above cunditio~~~ and the rotor head is trrtniPg, the pump will build up Iqdzsdic pressure in the system t&abat 200 psi, for stopping the rotor system. Five secaads afterthemtorheadstops a if the swikhispfaced awiththemtcn system stopped, h ~ p t e s s r a e i s ~ m a t i c i i l l y i n n e a s e d t o a t > a a t 4 3 0 p s i to prevent rotation of the rotor system when parked. An accumulator in the system will hold this pressure for aboat 8 hours. I£ -c pressr~e decreases while paxked, the pump w i l l be automatidly switched on to restore pressure in the system, provided the rotor brage sraitch is left ON, REL is a momenkry switch position which opens the control valve, rdeasing hykzulic pressare and the rotor brake. When the switch is released, it will mwe to OFF and the rotor brake system will be deenergized with power o£f. With p e r on, the rotor brake will release in the OFF position if the ROTOR BRAKE light, actioated by the 15 psi pressure switch,

Y!?

ism

NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Part 2 Section I Description

F U mMANUAL

Rotor Brake Cantion Light The ROTOR BRAKE caation light on the cacrtio&advisory partel (Figare 1-25) will be on when there is 15 psi or above hgdraulic pressare in the hydranlic system. The cadion light w i l l go on when the rator brake is applied, i d cfhaf the hydraulic pump is pressurizing the system. The light should go on before the brake puck begins to move. The light will be off when the rotor brake is off. The caation ligM circuit is connected to the dc essential bus by a circuit breaker marked RTFt BRK WARN. MANUAL ROTOR BRAKE SYSTEM Tfre rotor brake is a manually operated, hydradically-actuated system capable of shntdown against one engine or holding the totor against two engines during start and ground idle. The rotor brake system consists.of a master cylinder, pressure gage, accumdabr with relief =.be, rotor brake, pressure switch, and cadion light.

Master Cylinder The master cylinder (Figure I-9B)is mounted on the cockpit overhead. It is ~per&edmanually to provide hydradie pressure to ate the rotor brake on the main gear bas. The rotor brake is applied by pdlling down and ppshing forward an the master cy3ixier handle. Two successive strokes are required to develop the pressure necessary for sbppbg the rotoreffediveQ. Thelimerequbedto~tberatorsystemcanbecontrow by varying the amount of master cylinder handle applicalion &ring the second sboke. For normal operation, the totor brake should nut be applied mrtil rotor speed decreases Mow 65% Nr.

-

-nisThe aeamdabr and relid valve are maanted an t o p of the main i siw accessory gear k The accmndabr allows f o r ezqtansim uf the hydradic fluid, relieves spstem pressure if it goes over 300 psi, and a b sorbs pressure surges during brake srpplicafion The relief valve will by at pressures above 300 psi rator brake applicatiop. a

The rator brake receives hydm~Eower fram the master cylinder and acmmuhbr. It consists d two cylinder assemblies morrnted 180° apart, withtkebmkediscrat;ctirrrgthzorrghtkm. Thebrakedliscisboftedtothe main gear barrr t a i l bkeoff flange.

NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Part 2 Section I Description

NGHT MANUAL

Figure 1-93. Manaal Rotor B a k e Master Qlinder

ROTOR m

m

The rator configrnation is a single main rotor and anti-tome tail rotor. Both spstems are driven throagh the tansmission, a d blade angles are controlledtllxnl@the£li@~lspstem.

.

The main rotor cansists of main rator hub bolted to the main rotor drive sfiaft, fopt main zotor blades, blade dampers, a swashplate assembly, blades, and a bifikr vibration absorber. The bhdes are e h e d to the main rotor lxh bg elasbmeric bearings which permit tbe blades to flap vertically, hunt hotizontdlly, and mtate about their spanwise axis. At low rator speeds, anti-Oapping restrainers limit the upward movement of the blade and droop.stops limit the dawnward movement of the blade. Droop stops release at about 50%to 60%Nr. 'Hpdzrmfic dampers minimize hd5ng movement of the blades aad prevent shock to the blades when the zotor

NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Pan 2 Section I Description is started or stopped. Flight control inputs are uannnitted to the rotaring biades through the swashplate. The bifilar vibration absorber, mounted on top of the main rotor hub, absorbs 3-per-rev and 5-per-rev vibrauons usually transferred to the aircraft. Main Rotor Blades The four main rotor blades consist of a timiurn spar and a Nomex honeycomb core covered by a fiber glass skin. The leading edge of the biade is proteaed by titanium and nickel abraion strips. A swept-back tip improves performance and reduces noise generated at the tip. Grounding of metal fitting and shielding with aluminum mesh provides lightning protection.

TAIL ROTOR SYSTEM A cross-beam tail rotor blade system provides anti-torque action and directional control. The blades are of graphite and fiber @assconstruction. Blade flapand pitch change motion is provided by deflection of the flexible wdphite composite spar, eliminating all bearing and lubrication. The spar is a continuous member running from the attachment t of one blade. through the center hub. to the t of'the opposite blade. Like the main rotor blades. the four tail rotor blades have a fiber ass skin that is internally ed by a Nomex honeycomb. Flight control input is transmitted t o the blades through control horns that twist the spar. Grounding of metal fitting and shielding with aluminum mesh provides iightning protection-

FUEL SYSTEM (Figure 1- 10) Each engine has its own complete fuel system which under certain conditions can be rvnnetted t o the opposite engine by crossfeed vdlves. h c b system is a suctiontype supply system. and both systems operate in a similar manner. Two fuel tanks

NOVEMBER 21,1978 Revised March 2.1989

-

...~ ..-

FLIGHT MANUAL

Part 2 Section I Description

are mounted side by side beneath the baggage compartment floor. The tank on the lefr supplies fuel to the No. 1 engine and is gravity-fueled through a filler neck on the left side of the fuselage. The tank on the right supplies fuel to the No. 2 engine and is gravity-fueled t h r o u a a filler neck on the right side of the fuselage. Suction to draw fuel from the fuel tank to the engine fuel control unit is provided by the engine-driven fuel pump and jet inducer unit. Normal fuel flow is from the tank t h r o u a the DIR position of the fuel selector valve, through the jet inducer and filter assembly, and into the high pressure stage of the gear pump. Fuel at pump pressure is supplied to the g s producer fuel control. A by line returns fuel from the fuel control to the jet inducer and also to the primer line leading to the opposite engine. The fuel crossfeed system permits both engnes to operate from either fuel tank, or either engine to operate from the opposite tank. Fuel crossfeed is used by positioning the fuel system control handles as shown on Figure 1-1 1. The primer system consists of a primer line extending from the fuel control by of one engine to the crossfeed line leading to the opposite engine. A primer valve is contained in this line. One e n s n e must be operating to supply primins fuel to the opposite engine. Engine priming is used by positioning the fuel system control handles as shown on Figure 1-12. Xormally, engine priming is not needed, since check valves in the system prevent fuel line drainage back into the fuel tank. NOTE Extended running of an engine in DIRECT with the opposite engine's prime switch at OK and fuel lever at PRIME or XFEED can result in vent overflow if,tanks are full. FUEL LEVERS The NO. 1 ENG and NO. 2 ENG fuel levers are on the engine control quadrant (Figure 1-4). The levers are connected by mechanical linkage to the fuel selector valve in each main fuel line. The quadrant has four positions marked OFF-DIRPRIME-XFEED. Detents on the valve at all positions except PRIME indicate proper positioning of the lever. OFF closes the fuel line between a fuel tank and its respective engine but does not close the crossfeed line to the opposite fuel system. DIR opens the fuel line between a fuel tank and its respective engine. XFEED opens the fuel lines between this engine and the opposite tank. PRIME, although not a true valve position, is used if the fuel line leading to one engine has lost its prime. Fuel under pressure from the opposite engine or through the ground primer l the evacuated line. The fuel line must be full for the connection is used to f suction-type engine fuel pump to draw fuel from the fuel tank. Normally, priming is not necessary, except possibly after fuel line maintenance. In addition to manual positioning, the lever will also be moved to OFF when the respective T-handle is pulled to arm the engine fire extinguisher system.

NOVEMBER 21,1978 Revised March 2,1989

I

Pan 2 Section I Description

FUGHT 'MANUAL

rn

PRESS IND

NO. 2 FUEI PRESS IND

PRIME

ENGINE

I

GROUND PRIME CONNECTORS n n

NO. 2 ENGINE

FUEL QUANTITY

/I

NO. 1 FUEL TANK

C

Figure 1-10. Fuel System

NOVEMBER 21,1978 Revised September 7,1990

N0.2NEl TANK

I

Part 2 Section I Description

NCCn MANUAL

ENGINE PRlME SWITCHES

The NO. I EXG PRIME and NO. 2 ENG PNXE switches, on the overhead control (Figure 1-27) a r e lever-lock switches. Each switch is connected to an electrical primer valve in the primer line of the opposite engine. To prime an ergge, the switch is held to PRIME, with the fuel lever in PRIME. See Figare 1-12 for additional priming requirements. The No. 1 engine primer circuit is connected to the No. 2 dc primary bus by a circuit breaker marked FL ' TEL PRIME 1. The No. 2 en,* primer c m t is connected to the No. 1 dc primary bus by a circuit breaker marked FUEL

PIUME 2.

/

FUEL =SS REQUIRED

1

NO. 2ENG

OFF

NO. 1 TANKTO BOTB EN-S

NO. l E N G & NO. 2 ENG

DIR

NO. 2 TANKTO BOTH ENGINES

NO. l E N G & NO. 2 ENG

NO. 1 T A N K T O NO. 2 E N G m

NO. 1 FUEL

LEVER

-1

1 I

7

NO. 2 TANKTO

NO. 2FUEL LEVER

I

I

-=

Dm

NO. 1 ENG

NO. I ENGINE - --

-

--

Figare 1-11. Fuel Crossfeed Operation

FUEL PKESS

=Q=

NO. 1 FUEL LEVER

NO. 2

FUEL

PRZME

UVER

SWITCH

TOPRIME NO. I ENGINE

TOPRZME NO. 2 ENGINE

NO. 1ENG

PRllVIE NO. 1-

DIR

PEUME

NO. 2ENG

PRmE

Figure 1-l2, Engine Aime NOVEbfBER21, 1978

Revised M y 26, 1982

I

I I

I

FUEL QUANTTTY INDICATOR

-

1 . W indicator, marked MAIN FUEL LBS x 100, is a dual indicator that shows the fuel qpantity in each fxel tank. Fuel grrantitp data is ~~d by probes in the fuel tanks and transmitted electrically to the indicator. The indicaiing system is connected to the dc essential bus by circuit breakers marked NO. 1 FUEL QTY and NO. 2 FUEL QTY.

Ah

TOTAt CAPACITY

I

U.S. Wens

Imperial Gallans

Lites

U.S. Gallons

NO. 1

143.2

119.2

542

140.6

NO. 2

143.2

119.2

542

TOTAL

286.4

238.4

1084

Imperial Gallons

Liters

117.1

532

140.6

117.1

532

281.2

234.2

1064

'

Fuel Gage Test Button The FUEL GAGE TEST battan an the fuel gage test and digital light switch (Pigare 1-14) is ased to test the fuel quantity gages for proper opera!tion. With tic power available in the essential bus and the test Man pressed in, the fael quantity gage readings should decrease slowly to zero. When the button is released, the gages should reto the original fuel qpantities.

FUEL LOW LEVEL C A m O N LIGHTS The el F'UEL LOW and 62 FUEL LOW ca&ian lights on the ~001-advisory

I (Figme 1-25) light when 85 to 100 pocmds of fuel remains in the correspanding tank The lights are connected to the No. 2 dc primary bus try circuit breakers marked NO, 1 FUEL LOW WARN and NO. 2 FUEL LOW

WARN.

NOVEMBER21, 1978 Revised Jdip 26, 1982

Pan 2, Section I Description

FLIGHT MANUAL

ELECTRICAL SYSTEM The primary source of elecmcal power is a 28 vdc (direct current) system. A I 15 vac (alternating current) system is powered by an ac generator and one standby invener.

I(

-

ACCEL FUEL V A L E !XOW

FIRE

TS]

EX7

, ACCU

FUEL VALVE FLOW

NR

FIRE ECT

TS

N1

EAPS W PRESS

N2

TOR0

BIAS

COW

HYD

XMSN

HOOK

S m DET

LT INST DIM

TIR GYRO LT

WLT

ADF

VOR LT

? W . X W

I\HPL

FUELLO /N0.1Eb rHY001LULSN7rWARN -\ FUEL PRESS TEMP NO. 1 N0.2 PRIME

7LTG OIM PLTINST CSL

PgS

PLT

CABIN

7FLOAT 2 PWR 1

-

PWR

FLT

Am -BY

ADF FLT

N0.2 XPDR

FLOOD FONE

r SEC LOG LT-\ CONT

WAW

1

I

-LO

RDNG LTS

WTHR

SWG HTRCONT RDR

AD1 WLT

FD

LT

Figre 1-15. Upper Circuit Breaker Panei, DC Primary Buses Typical AC Generator System (Aircraft Serial Numbers 760270 and Subsequent) NOVEMBER 21, 1978 Revised May 15, 1992

I

Pan 2, Section I Description

FLIGHT MANUAL

-

-

Circuit breakers in the pilots' compamnent are on two s (Fie- 1- 15, 1 16, I 1i , . and 1-18) on the flight control housing between the pilot's and copilot's seats. The upper contains circuit breakers for electrical circuits connected to the No. 1 and No. 2 dc primary buses. The lower contains circuit breakers for electrical circuits connected to the dc essential and the KO. 1 and No. 2 ac buses. The following circuit breakers are on the dc junction box in the electrical compartment; GEN CONT NO. 1 and NO. 2, BATT CONT, VENT BLOWER PWR, LNV PWR and ROTOR BRAKE - PWR and CONT.'

Fizure 1-16. Upper Circuit Breaker , DC Primary Buses Typical AC Generator System (Aircraft Serial Numbers 760001 through 760269) NOVEMBER 21, 1978 Revised May 15, 1992

-

Part 2 Section I Description

DC POWER SUPPLY SYSTEM

The primary power source for the dc system is two starter-generators. The secondary power source is the battery. For g m d operation, an external source of dc power may be used to supply power to the electrical system.

NOVEMBER 21, 1978 Revised FEBRUARY 22, 1985

Part 2, Section I Description

FUGHT MANUAL /

DC GEN INV AC GEN RESET WARN WARN

NO. 2

FUEL

7FIRE D m C T O R ?

QTY

N0.2

NO.l

M

NO. 1

FUEL QTY

CHIP

DE:

\

DC G W STR RESET CNTOR

"RE E*

COW

] N

2

1 SPD OIL PRESS Nl

X16N OU.

RTR BRK /-WARN WARN

PRESS

NR

r S V O PRESS7 CLlV TRIM 2-WARU1

TRIM

WARS

N

r

CHIP D E T l T i R GYRO MGB I i T G S PILOT

I

I

cA0112 SI

Figure 1-17. Lower Circuit Breaker , DC Essential and AC Buses - Typical AC Generator System (Aircraft Serial Numbers 760270 and Subsequent)

NOVEMBER

21, 1978 Revised May 19, 1992

1~27

Part 2, Section I Description

DCGPl

k

WCS

I PRIAC

-.

. ~ -~

.

.. -

FLIGHT MANUAL

INV

WARN

ACGD4

W0.2

-RREDElEETOR-

WARN N E l O T Y N0.2

.'

PLT

A n ' PLT

VOR PLT

CaPm

GYRO

O C W

WD.1

NO. 1

TEST NELOTY RESET

Arr LT

GYRO

RMI

corn

LT

FID

'

R 1

< PLT AD1

FID

LT

AD1

VOR

LT

2

-

2 ENG -1 I N U T -ICE

Figure 1-18. Lower Circuit Breaker , DC Essential and AC Buses - Typical AC Generator System (Aircraft Serial Numbers 760001 through 760269)

NOVEMBER 2 1 , 1-28

1978

Revised May 15, 1992

1

NGHT MANUAL

Part 2 Section I Description

No. 1 and No. 2 Starter-geneators A 200 amp starter-generator is maanted on the accessory gear box section of each engine. They frmction as engine starters when provided with dc power from the battery o r an external power source. For sfarter operation, refer to paragraph tilied ENGJNE STARTER AND IGNITION SYSTEM, in this section. After an engine start when tke engine is operating at idle, they function as dc generators. Generator control s regulate generator oatput and protect against wervoltage, undervoltage, reverse c m n t , and ground fanlt conditions. Generator Switches TFPo switches on the master switch p a w l marked GENERATORS NO. 1 DC and NO. 2 DC bave positions marked OR, OFF, and RESET. ON connects each generator to its power distribution system when the generator contml senses tbat generator output is within limits. OFF switches off the generator and disconnects it from the m n system. I€.the generator h a s been dropped aff the Iine due to a momentary overvoltage o r other fault ($1 or +2 DC GEN caution light on), placing the switch to RESET and then ON will restore generator operati& The reset circuits a r e connected to the dc essential bus by circuit breakers marked DC GEN RESET.

-

Generator Test Switches The CtEN 1and GEN 2 test switches are on a small marked =ST (Figure 1-20) on the copilot's side of the instrument , Each switch has three positions, marked GM) FAULT, OFF, and OVERVOLT. GND FAULT and OVERVOLT are nsed to test that the generator w i l l drop off the line if either of these conditions occur. To test the system, the snitch is held at either GND FAULT o r OVERVOLT for 5 seconds, with the N1 of each engine at ID= o r above. The generator should drop off the line as of the generator cardion light. To restore geneaindicated by the tor opethe g e n e m r switch is placed to BESET and then ON. Generator Caution Ljpfitc: The #l I X GEN and #2 DC GEN camtian lights, an the caulion-advisory , o r the generator go on when the respective generator has switch is OFF and the dc spstemis opexafiag fmm anathgr power source (ather generator, battery, o r external power). The lights are connected to the No. 1and No. 2 dc primary buses by breakers marked DC GEN WARN.

DC Generator Hot Caution The fl DC GEN HOT and #2 DC GEN EIOT czmtion lights (aptiondl), on the caatiobadvisory , w i l l go on when atcessive generator exbatst duct temperatures a r e sensed, indicating a geeerator overtemperatme condition. The lights a r e connected to the No. 1and No. 2 dc primary buses by circuit breakers marked DC GEN WARN.

Battery A nickel-chm I? amp hour o r 34 a m p hwr ( o p t i d ) b&terg is installed in the electrical comp.Hzm& The battery is used for limited w m d operation and as a secadary source of puwer in Bight, The battery provides power only to the most esseniial equipment. A battery averhqteratrae protection system auhxmiically disconnects the battery from the electrical s g s t e m i f m r p t e m p e a t a r e , duetonormalreckrgmg, reaches63OC NOVEMBER 21, 1978 Revised July 26, 1982

1-29

I

-

NO. 2 OC

NO. 2 DC STARTER0ENERATOR AND GENERATOR CONTROL

-

I RLSEl

NO. 2 OENERATOR CON

-

C\IC

NO. 1 DC STARTEROENERATOR AND (XENERATOR CONTROL

8;f3 3

A

M

BATTERY C O N 1ACTOR

CAUTION LldHT

NO. 2 ' D C PRI BUS

EXTERNAL POWER

ENTERNAL POWER OR

I

BUS TI1

UI; tss

BUS

CAUTION LIOHT

GENERATORORS 1. CONNECT EACH GENERATOR TO ITS PRIMARY BUS I F GENERATOR OUTPUT IS ACCEPTABLE TO GENERATOR CONTROL . 2. CONNECT EITHER GENERATOR TO BOTH PRIMARY BUSES IF OTIIER GENERATOR IS OFF OR HAS FAILED.

BATTERY OR 1. CONNECTS BATTERY TO DC ESSENTIAL B U S ONLY. DISCONNECTS BATTERY FROM DC ESSEIWAL BUS WHEN EXTERNAL POWER IS I N USE OR WHEN BATTERY SWITCH IS OFF.

EXTERNAL POWER OR 1. CONNECTS EXTERNAL POWER TO DC SYSTEM EXCEPT WHEN ONE OR BOTH GENERATORS ARE OPERATING.

BUS TIE OR 1. CONNECTS DC ESSENTIAL B U S TO NO. 1 DC PRIMARY B U S WHEN ONE OR BOTH GENERATORS ARE OPERATING OR WHEN EXTERNAL POWER IS I N USE. 2. ISOLATES BATTERY OUTPUT FROM PRIMARY OC BUSES W I K N OPERATING ON BATTERY ONLY.

GENERATOR CAUTION LIGHTS 1. ON WHEN GENERATOR IS OFF OR INOPERATIVE AND DC SYSTEM IS POWERED BY OTHER GENERATOR. EXTERNAL POWER OR BATTERY.

BATT HOT WARNING LIGHT 1. ON DURING PERIOD OF BATTERY OVERTEMP. WHEN O N BATTERY IS DISCONNECTED FROM ESSENTIAL BUS.

BATT OFF CAUTION LIGHT 1. O N WHEN DC SYSTEM IS POWERED AND BATTERY SWITCH IS OFF. EXTERNAL POWER IS I N USE, DURING PERIOD OF BATTERY OVERTEMPERATURE OR OTHER BATTERY MALFUNCTION.

BUS TIE OPEN CAUTION LIGHT 1. O N IF GENERATOR OUTPUT DOES NOT GO THROUGH THE BUS TIE OR TO POWER THE DC ESSENTIAL BUS. WHEN ON. DC ESSENTIAL BUS IS POWERED FROM BATTERY ONLY.

Part 2, Section I

Description

TEST AC GEN

1

1

FAULT

FAULT

OVER VOLT

I

I

UNDER

~5 F

OVER VOLT

FEEDER FAULT

F

Figure 1-20. AC and DC Generator Test s (145°F). The battery is automaticdly reconnected to the elecuical system when its temperature drops to 52OC (125°F) or if both generators malfunction. The battery will also be disconnected from the electrical system when external power is applied. Banery Switch The BATT switch on the master switch p q e i (Figure 1-9) has positions marked ON and OFF. ON connects the battery to the dc essential bus except when dc external power is in use. OFF disconnects the banery from the dc essential bus. Banery Hot Warning Light The red BATT HOT warning light on the cautionladvisory (Figure 1-25) lights whenever the battery is disconnected from the electrical system due to overtemp. When the battery cools it is automatically reconnected to the electrical system and the light will go off. The warning light is inoperative when the banery switch is OFF. The BATT OFF caution light on the caution-advisory will aiso light when the BATT HOT light goes on. Battery Off Caution Light The BATT OFF caution light on the cautionladvisory (Figure 1-25) lights whenever the dc system is operating and the battery is disconnected from the dc essential bus. The light is connected to the No. 1 dc primary bus by a circuit breaker marked BATT-WARN. External Power A 28 vdc external power receptacle is on the right side of the fuselage above the main landing gear door. External power will supply power to the entire NOVEMBER 2 1, 1978 Revised May 15, 1992

Pazt 2 Section I

FUGHt MANUAL

Description

dc system. When external power is being used, the battery is automatically disconnected from the dc spstem. When either generator is operating, exterml power is automatically discounected from the dc system.

The EXT PWR switch on the master switch (Figme 1-9) has two positions marked ON and OFF. OK connects dc external power to the dc mtem except when one o r both dc generators are operating. OFF disconnects the external power supply. DC POWER DISTRIBUTION SYSTEM DC power is distriboted through three buses. The No. 1 dc primary bus is powered by the No. 1 generator and the No. 2 dc bus is powered by the No. 2 generator. The dc essential bus is connected to No. 1 generator outpnt though a bas tie or. If either generator is inopedive as indicated by the lighting of the appropriate DC GEN cantion light, the other generator w i l l provide power to all three buses. If bath generators a r e ir+ operative and external power is not in use, only the dc essential bus will be powered from tbe hattery. .Thiscondition is indicated by the EgMing of bath DC GEN and BUS TIE OPEN caution lights. Ekternal power will power all three bases unless one or more generators a r e opeating, &en external power will be disconnected from the dc system. Bus Tie Open -tion

Light

The BUS TIE OPEN caution light on the canlion-advisory ( m e 1-25 EgMs when the circuit is open between tbe generators o r the extetnal power receptacle and the dc essential bas. The& only the battery can provide power to the dc essential bas, and the battery wiLl not be recharged by generatot output. The canlion light is connected to the dc essential bus by a circnit breaker marked BUS TIE WARN.

The bus tie or will sometimes close momenbrily when some dc equipment is switched on with the battery as the only dc p e r -e. This transient will be ip dicated by a flash of tbe BUS TIE OPEN caution Light and possible £lashing of the DOOR OPEK o r other caution lights and a flicker of those ~ e n powered b by the bus. dc

NOVEMBER 21, 1978 Revised JUNE 28, 1984

Part 2 Section I Description

AC

-

NGHT MANUAL

POWER SUPPLY SYSTEM

(AC G-TOR

- ( ~ i g a r e1-22)

The primary power source for the ac electrical system is a generator, and the secondary power source is an inverter. For grorrnd operation using external dc power, the inverter may be used as a source of ac power.

A three-pbse 7.5 KVA generator is mounted on and driven by the main gear box. A generator control unit regulates generator output and protects

against overvoltage, undervoltage, under frequency and feeder fault. The generator conbml circuits a r e puwered by a dc permanent magnet generator (part of ac generator). Three-phase ac is required if the windshield heater

is installed. AC Generator Switch

-

The GENERATORS AC switch on the master switch (Figure 1-9) has three positions, ON, OFF, and TEST. ON connects the generator to the ac power d i d z h h o n system when the geneator conh.ol senses that generator output is within limits. O F F switches off the generator and disconnects it from the ac sys&em. If tbe generator is dropped off the line due to a momentary overvoltage or ather fa& (AC GEN cantion light on), placing the switch to OFF then ON ppill restore generator opeation. The TEST position may be used to test genexator output before comrectiag it to the ac c k s k b d i o n system (switch on). If genetator o d p d is , placing& ~ h t o ~ ~ ~ ' a h e n o p e r a i 1j a0 g0a% i w~i ~ llputo9~the~~ ~ c a a t i o n ~

R

NOVEMBER 21, 1918 M FEBRUARY 22, 1985

Part2seCfinnI Description

AC GENERATOR AND CONTROL UNK

F U M MANUAL

wo. IAC

:& F

tEn

I-[ CAUTION U G H T PRl BUS OR ADVISORY LIGHT

4

NO. 2 DC PRI BUS

CAUTlON U G H T

WMRIU1

INVERTER

P R I M BUS COWrACTOR 1. CONNECTS AC GENERATOR TO ALL BUSES AFIER MAIN TRANSMfSSlON IS OPERATING AND GENERATOR OUTPUT lS AC#VTABLE TO GENERATOR CONTROL UNIT. ,2 CONNECTS INVERTER TO AC PRIMARY BUSES 1F AC GENERATOR IS INOPERATIVEOR HAS FAILED. AND THE INVERTER SWITCH IS IN THE ON POSITION.

AC 6EN CAUTION UCHT 1. WHEN OK INDICATES THAT AC GENERATOR IS lNOPERAlWE OR HAS FAILED.

INVERTERCAUnONUGM 1. WHEN ON. INDICATES THERE IS NO INVERTER OUTPUT.

INV ON ADVISORY U6HT 1. WHEN OKINDICATES THAT INVERTER IS SUPPLYING POWER TO AC PRIMARY BUSES.

Figme 1-22. AC Electrical

a-

- 'Ilspical AC

AC Generator Test Switch A . ~ ~ J I ~ ~ ~ u ~ ~ ~ T(Figure E S 1-20)is ~ A C onthe G E mpiht's N sideof the instrmnentpamd, T h e p a n e l f r a s t w o ~ , o n e ~ V N D ~ O L T OFF and OVEaVOLT and the second marked FEEDEB FAULT and OFF. Theswitches~used~totestthat~geneafor~dfopaflthelineifarry o f t h e a b o P e ~ E ~ a ~ c. b c o n ~ i s t e s t e d ~ b g ~

-

.

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aswitchatthedesiredpsitionforabout58e~ondsandthearekasingit. The generat~rshodlddro~offthelbeas ~catedbythe GENcanlionlight, Torestoregenerator.aperation, WAC-i s p k c e d t o O F F , thenm.

ofthe AC

~2SectionI Description

The AC GEN caution light, on the caution-advisory (Figare 1-25) W when the dc essential bas is energized and the ac geneator is inoperafive, not c m e c t e d to the ac distribrrtion system, o r bas failed. The caation

-

iscannectedtotheNo. 2 drimarybusaadthe dcessentialbrrsbycirtrrit breakers marked AC GEN WARh7.

AC Generator Bearing Caution Light The rotor shaft of the ac generator has a set of main bearings and a set of

atrxiLiarp bearings. Normally the main bearings carry the rotor shaft load and the &iarg bearings are clear of the shaft. If the main fail, the a n x i k r y bearings will take over the rator shaft load. Then, a bearing failure switch w i l l light the AC GEN E A R l N G light an the caution-advisory b e a r i q p is limited to 5 (Figure 1-25). Operation on the h o ~ r s . The cantion light is connected to the No, 2 dc 'JP -I bas and the dc essentidl bus by circuit breakers marked AC GEN WARN.

The secondary power s w r c e for the ac electrical spstem is a 600 VA inverter. !RE inverter is powered from the No. 2 dc primarg & a circait breaker marked INV PWR an the dc junction box. For ground operatians, the inverter may be used as a source of ac power udil the rotor system is engaged and the ac generator is driven by the main &ar box.

laverter oaipxt is c o M e d by an INVERTER ON-OFF switch on the master swftch ( F i 1-9).

+

h r t e r Fail Light SomehelicaptershaveanINFrEILTEBcaatianlighton~~gP,-~ (Figure 1-25) that lights when the inverter fails while either supplying p e r to the ac system o r acting in stand-by as a backup to the ac gerterator.

I

AC IWWER O D N -

(AC GENERATOR

NOVEhdBER 21, 1978 Revised July 10, 1987

Part 2 Section I Descri~tion -+

FLIGHT MANUAL

FLIGHT CONTROL SYSTEM ~onventionalhelicopter flight controls consist of a collective pitch lever a a cyclic control stick to control main rator blade angles, and fail rotor pedals to control tail rotor blade angles. A *stage servo system reacts to loads imposed on the rotor sgstem and reduces control forces required by the pilot. Collective and cyclic trim and a force gradient system permits t. - g of the controls in the cockpit to the desired position. A set of tiad controls for copilot m a y be installed as optional equipment. MAW ROTOR FLIGHT CONTROLS

Control movements from the collective pitch lever for vertical coatrol aPd from the cyclic control stick for directional control are transmitted by mechanical linkage to a mixing unit which combines the lmo inputs. The combined input is then transmitted to the stationary ring of the sarsslpiate by mechanical -e, and through the *stage hydraulic s e m -em. Movement is tansmitted from the stationary to the rotating ring of the swashplate to varg the pitch of the mnin rotor blades. A collective to p w coupling aatomatically changes kil rotor blade angle and tln-ust to compensate for changes in main rotor torque when collective pitch is increased or decreased.

Collective Pitch Lever The collective pitch lever (Figure 1-5) increases or decreases the collectipitch of the main rotor blades. The collective pitch lwer may be trimmed to any desired position by use of the trim system. Collective pitch lever movement from the trimmed position is resisted by a gradient spring.

Collective Pitch Lever Trim and Force Gradient Spring

I

The collective pitch lever may be trimmed to any desired positiun and held there with a ma$p&ic bake. The trim system operates from the dc essential bus through a circuit breaker marked CLTV TRfM on the lower circuit breaker . The trim is cantrolled by a master STICK I?UM CLTV switch on the master mtch with positions m k e d ON and OFF and a switch on the collective pitch grip =Bed TRWI REL. When the magnetic brake is energized electrically, i t holds the lwer positio~ When the magnetic brake is de-energized,the lever is released, With the master switch ON, collective lever may be trimmed by pressing the pdbutbn switch to de-energize the brake, moving the 1eve.r to the desired position, and then releasing the pnshbrr#on. The lever will remain trimmed to t?~& psi-% Any movement fmm the trimmed position w i l l be resisted by the force gadient spriPg which creates a stick 'Yeel''.

*

NOVEMBER 21, 1978 Revised March 17, 1986

cyclic control Stick The cyclic control stick (Figure 1-5) changes the pitch of the main rotor blades to provide longitudinal and Meal control. The cyclic control stick may be trimmed to any desired posilion by use of the magnetic bakes. Cyclic ctx&ol stick movement from the trimmed position will be resisted by a gradient spring.

Cyclic Control Stick Trim and Force Gradient Spring The cyclic control stick trim system fxmctions similar to the collective pitch lever, except that two magnetic brakes, one for pitch control, and one for roll control, are used The trim system operates fromthe dc essential bus thmogh a circuit breaker marked CYC TRIM. The master switch on the master switch is marked STICK TRIM-CYCIJCand has positions marked ON and OFF. The pushbatton switch is on the cyclic stick grip and is marked TRIM REL. Pit& Bias Actuator

1

Some helicopters are equ@ped with a pitch bias -r that improves hgitudinal stabilitp by compensating for changes in frrselage attiimde caused by changes to airspeed or power. An airspeed transdacer senses changes in airspeed and a collective pitch transciucer senses changes in collective pitch lever position. These a e s are combined and transmitted efectri. cally to a pitch bias actuator, The actoator applies corrective inprrt to the : Elight cantrol spstgm wilenever changes to airspeed (cyclic) or pow= (collective) are made by the pilot The pitch bias actuator is limited to tD 11% forward;md17%aftoffdll~~inprrtofthefiightccmtro1system.The corrective inprrts result in a constant fodisplacement ad the cyclic con-1 stick as airspeed is increased and e c e s changes in fuselage attitudewithcollectivepikhdxmge6. Thesystemis0peatiandlwhentheai.rspeed mmshcer senses aborrt 45 knots airspeed or above, Electrical power is6qpliedtothesystemframtkeNo. 1 p r j ~ ~ ' d c b P s ~ a c i r c P i t breaker marked BlAS ACTR. Channel Monitor Switch

NOVEMBER 21, 1978 Reviswl JUN3 28, 1984

WCHf MANUAL about 45 hots airspeed, the indicator pointers s h d be full left. Above

a b d 45 h o t s , the NO. 1pointer should follow movemeats of the NO. 2 pointer. The AFCS indicator is installed an all helicopters even if AFCS is not installed.

TAIL ROTOR FLIGHT CONTROLS Tail Rotor Pedals The tail rotor pedals coatrol the pitch and thnrst of the tail rotor bbdes to the compensate for main rotor torqae and to provide a means of -c heading of the helicopter. The pedals have --operated wheel brake pedals. Cantrol rods and cables connect the pedals to the pitch rjlarvrinP mechaninm at the tail rotor. Caatrol forces are relieved by the two-stage tail-rotor servo. A pedal clamper gradient springprovides Yeel" in the tail rotor pedals by resisting pedal movements, to prevent overcontrol. Tail Rotor Pedal Adjustment fLnobs At a i l rotor pedal adjustment knob is belm the ' ' . ent and centered between the pedals. The knob is marked PEDAL ADJUST-TURK. Arrows indicate direction to tum for FWD and AFT adjustments.

Mechanical coUective b yaw coupling incorporated in the flight 1 system provides automatic tail rotor pitch (thmst) changes proportional b collective pitch (torque changes). The cazpling reduces pilot workload by abmaiically cmqms&ag for main rotor brqne as collective pitch is increased or decreased. Thcmgh mechanical stops prevent tail rotor pitch limits from being exceeded, mechapical tan be felt m the carrtrols at certain control extremes. F 'down collective and apdappicatim offull right pedal w i l l result in upward movement of the collective. Ruther movement d the right pedal, however, prodaces no further change in iail rotor pitch, since the tailrotor servo is at its mecbanica;l stop. I&ewise, at high collective settings, application of left pedal may re& in system limits being sttained before the pedal sLops are reached. Frvther application of left pedal will result in downward movement of collective and no change in tailrotor pitch. The extreme control positions where the cdllective to yaw envefee&ack occurs are not normally encoantered within the normal lope but may be felt during dynamic maneuvers; e. g., arresting a rapid descent rate in a right crosswind at maximum gross weight.

SELF-C E N T E m G TAlL ROTOR CONTROL QUADRANT The self-centering aft quahnt will retarn the tail rotor blade # k h to -2" in case of a dual control cable failxre or single faibPre rmapposed by the temainirrg cable. The qoadant uses two pivoting atms with concerrtriemounted torsion springs at the base end and reaction rollers at the outer end. When c o d m l cable tension is lost, the respective torsion spring pivots its arm to a fixed reattion plate imparting a r o t a t i d force to tbet -4 Inthe case ofthe single cable failure, tbisforce,may be reacted bythe pilot though the remaining cable b retain a desired trim condifion&pendent Prefaibrrre pedal position aad failed cable combination. In the case ofthe dualcable failure, t h e ~ w i l l r o t a t e r m a p p o s e d t o t h e - 2 °setting.

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1-40

NDVEBdBER 21. 1978 Revised JUNE 29, 1984

Part 2 Section 1 Description NGHT MANUAL

The -2" setting is a compromise that rein coordinated level fligM at about 40and 1% EaAS in addifion to varied combidions of trimmed and rmtrimmed climbs, descents, a x h v b t i v e descents, and level flight conditio~ ap though VNE. Flight conditions at airspeeds &SS than 20 lntots may be considered outside a recoverable envelope using -2'.

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A single cable failure affecting rigfit cable will retain normal lefk yaw contml from the -2" setling to full left peQL In level the helicopter will yaw left beween 40 and 135 gZAS reright lateral cyclic to mainkin track A left cable failme will retain normal right yaw control from the -2" setting to full right pedal. The helicopter will trim normally in level w a t a i r s p e e d s . 4 0 - 135KLASbirt~ri.gbtatairspeedsoatsidetbat band. Left cyclic inpat will - ' track chy.ing y m conditions. Druiag single cable failure conditions nonnal collective to yaw coupling will be available through the remaining cable and its limited control range. The respective p e w force of the remaining cable varies from near zero at -2" to as much as 20 pounds at the control extremes. '

Eeficapter response to a cable (or cables) faiftve WU depend Upon airspeed and flight condition (level, climb, or descent) at the thee af fairare.

Z)oalserw,~inthe~aadatandemseroo&inthetailrotor~ control systems react flight loads ftom the rotor system and r e b e the force required to operate the f&bt corttrols. The dual s e m amits consist af a first stage and a second stage eachaf which operate on 3000 psi hpctraalic pressure fmm separate hydmdic pumps. The main gear bax. Full pressure is avaibble at ahout 10 &. Three main rotor dual serpounits are mounted onthe maingearbaxandattackdtothe stationary swashplate, The tail rotor tandem servo is marmted on the tail rator gear box. Eychnlic pressure for extenand the landing g e a x i s ~ b y t h e s e c o n d s t a g e h p d r a p l i c ~T. h e ~ a n d s e c o n d stage servo systems are controlled electrically by switches on the collective stick grips. Electrical irrterlocks between the &st and second sktge sys~preventanespstemfrombeingsbrrrt~~~is1600psiinthe apposite system. If one stage is shut ofl and a pressare loss occurs in the systeminoperation, theipterlockswjllswikkanthespstemthatwas

Tye--

origindllydmtaff.

The first and second sbge servo s y h s are cantrolled by a three-po* SERVO switch an the collectme pitch (Figure 1-51, The switch bas

NOVEMBER 21, 1978

Revised JUNE 29, 1984

Part 2 Section I Description

positions marked NO. 1OFF and NO. 2 OFF. The switch is normally centered with both stages operating. To tma off either stage, the switch is placed to the, correspanding OFF position. If there is a t least 1600 psi hydradic pressure in the other stage, the selected sgskm u5ll go off. The shutoff system is connected to the dc essential bPS by circuit breakers marked SVO SWTOFF NO. 1 NO. 2, When drral fLigM canSols a r e installed, the copilot's collectme grip bas a simibr switch. Either switch may be used t o shut off one dage; hmever, the same switch mast be used to turn that stage back on again. When one dage has been shut off, the other stage cannot be shut off by the opposite svktch.

-

-

NOTE If a dc electrical system failure has occurred, a malfrmctioning servo system m o t be shut of£. If a mallPnctioniag servo system has been shut off, it will be reactivated if a dc failure occurs.

FUGRT CONTROL SERVO HTDRAULIC PRESSURE INDICATOR A dual indicator marked HYD PRESS PSI X 1000 has scales marked 1and2 for first and second stage hydanlic pressure, The indicator is conuected to the No. 1aud No. 2 dc primary buses by two circuit breakers marked 3YD OIL PRESS.

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A bl SERVO SYSTEM and a Q SERVO SYSTEM caution lights, on the caution advisory (Figure 1-25) go on w h a me h m u l i c pressure in the corresponding stage dmps t6 1600 psi. The light will go off when pressure increases to 2200 psi. The same caution ligtcis atso sense a jam in one o r more of the s e m mi&- When a s e m becomes jammed, the corresponding caotion light will go M. Duriag the caution light wiU be kept on through a holding circrrit even if the w o n of the1-j pilot valve retprns to normal. The holding relay operates the landing gear interlock and therefore w i l l only function when the helicopter is airborne. To differerrtiate between a system pressure loss o r a jammed s e m , note the pressure on the system hydraulic pressrae indicator of the afcected system. A loss in indicated pressrae with the light on indicates a

loss in system pressure. No& indicated pressure with the caution @hi on could indicate a jammed s e m . The d o n light tircaits are connected to the dc esseatidl bps by circrrit breakers marked SVO PRESS-2 WARN 1 and SVO JAM-2 WARN 1.

A d u a l ~ A F C S p r o v i d e s s g b Z l i t p a b o u t t h e p i t c h ,~ a n d y a a r a x i s . The~sareredu&&, eachwithasepaateelectritatpowersmrce, vertical gpro, yaw rate gyro, finear actuators (pitch, roll, and yaw), and a control . Control aathority uf each cbannel is limit#l to !%, WWI a wmbined limited authority of 10%. I)ae to this Iimihd a d ~ ~ 3 the @ ,pilot may easily override AFCS iaprrts through normal of the fligW cmtroLs if an AFCS system should malfunctian. The gyros in each ckumel transmit signals to an amplifier which senses any deviation from tke iximmed attitado and headiag of the helicopter as select& by the pilot. Whenever there is

NOVEMBER 21, 1978 Revised JUNE 28, 1984

I

Part 2 Section I Descri~tion

F U M MANUAL

AFCS CONTROL AND INDICATOR

AfCS Phase 11 Control and m

t Breaker

s

any deviatkm, the amplifier transmits an electrical error signal to the lineat actuators, which are an integral part of the flight control system. ~Lineara~sin~~#arectiveactimtothetlig1lcantrol system. Airspeed switches in the copilot's pitot static system increase amplifier gain below speeds of 40 or 60 knots depending on the airspeed switch instdlled. This provides greater sensitivity to the AFCS input to the night controls, resulting in greater stability at lower airspeeds and in a hover.

I

NOVEMHER 21. 1978 Revised NZARCEI 2 4 1981

NGHT MANUAL

ATTITUDE RETENTION. Pitch and roll sa r e pnmided to each channel by a vertical gym. The vertical g y m is referenced to the horizon; however, the attitude afthe helicopter may be trimmed to an attitude desired by the pilot (e.g. attitude for desired cruise) which deviates from the horizontal attitude. .Once the helicopter is in the trimmed attitude the vertical gpro signal is nulled in the amplifier and there is no electrical i a p t to the pitch or roll hear actuators. When the attitude of the helicopter deviates horn the trimmed attitude, an e r r o r signal proportional to the deviation is sensed in the amplifier, which ttansmits an electrical signal to the linear actuators. The pitch and roll linear actuators i n the main rotor cyclic pitch flight control system apply a corrective input to retuxn the W c o p t e r to the trimmed attitude.

Helicopter attitude may be established o r changed by either of these methods: (1) The cyclic stick position may be retrimmed using the magwtic. brake. When the TEUM REL mitch is depressed, the e r r o r signal to the actrrators is nulled and the magnetic brake released. The cyclic stick is tka moved t o the desired position and the TRIM REL switch released. AFCS w i l l hold the newly established attitude. (2) The four-way beeper trim switch on fbe cyclic stick grip is normally used t o make small changes to helicopter attitude without cbang&g the cyclic stick position. A.essing the switch to any af the four positions : FWD, AFT, L, o r R, drives the actuators to produce the desired helicopter . attitnde. When the switch is released (centered), AFCS wiIl W d this atatrode. On some AFCS Phase I1 helicopters, the beep circuit is only activated for 3 seconds. If additional correction is necessary, release the switch momentatily and then press again to the desired position.

Yaw signals are ptwided to each ctannel by a yaw xaie ggro. When the hexicopter develops a yawing rate, an error signal proportional tothe rate is sensed by the rate gym in the amplifier, which transmits an electrical signal to the yaw linear actuator. The yaw linear actpator in the tail rotor nipm control system applies a corrective inprd to arrest the yaw rate.

AFCS COhmOL

Two identicdl AFCS CONTROL s (Figure 1-23)aze on the center cansole. 'RE forward controls the No. 1channel and the aft controls Each has three switches marked PITCH, ROLL, the No, 2 c-1.

NOVEMBER 21. 1978 Revised MAR- 24, 1981

1

Part 2 Section I Description

FLIGHT MANUAL

and YAW. Each switch has three marked positions ENG (en-ge), OFT, and DISENG (disengage). A light next to each switch lever marked NO GO will go on if there is an electrical malfunction in the system o r in the @n control. The NO GO lights will also be tested (go on)when the caution-advisory test button is pressed. m e ENGAGE position of each switch engages the AFCS for the respective axis. The OFF position of each m t c h centers the corresponding actuator and then removes power from the system. The lever- lock DISENG position of each switch disenpges the system with ihe actuator remaining in the position a t which the system was d i s e w e d .

AFCS INDICATOR The AFCS INDICATOR (Figme 1-23) is next to the A f CS control s on the center console. The contains a channel monitor switch with marked positions OFT, PITCH, ROLL, YAW, and PBA, and two indicators marked CH 1and CH 2. By selecting PITCH, ROLL, o r YAW on ~e rotary switch, the position of the respective actuators (LFT/FWD or RT/AFT) can be monitored inthe windows. Normally the indicator pointer should move about the center position. The PBA position is used to monitor the pitch r this bias actuator system. (Refer to paragraph tifled Pitch Bias ~ c t n a t t iin section. ) AFCS CIRCUIT BREAKERS (Figure 1-23) m e No. 1AFCS channel is connected to the dc essential Isus and the No. 1 a c primary bus by c i e t breakers marked AFCS DC ESNTL and AFCS PFU AC, respectively. Tke No. 2 AFCS channel is connected to the No. 2 dc primary bus and the No. 2 ac primary bus by circuit breakers marked AFCS from tttese circmt breakers PFU DC and AFCS PRI AC. Power is -mitt& to additional AFCS circoit breakers on a on the cockpit ceiling marked AFCS. The vertical gmos for the No. 1and No. 2 AFCS channels a r e connected to the No. 1a c primary bus by circuit breakers marked GYRO LT and GYRO P L I , respectively. The turn rate ggro for the No. 1AFCS &annel is connetted to the No. 1dc primary bus by a circuit breaker marked T h t GYRO LT. Tke turn a t e gym for the No. 2 AFCS channel is connected to the dc essential bus by a cirit breaker nrarked T/R GYaO PILOT.

AFCS Pbse m combines the limited authority capabilities of AFCS Phase II with automatic cyclic stick and tail rotor pedal positioning, heading hold, and automatic coordinated trva Pbase II is a dnal "inner loop" system with a combined limit of 40% control authority which imparts no movement to the cockpit flight controls. Pbase IU combines "inner loop" with "outer loop" which increases flight control authority to 100%. The "inner loop" senses small attibade and yaw rate deviations from those desired by the pilot, and automatidly applies corrections within its limited aufbority with no movement to the cockpit flight controls. For large inner loop deviations thel'aater loop" w i l l automatically reposition the cockpit flight controls to correct for the deviation. The pilot may w e m d e any AFCS inppts'thmq3 normal use of the night controls. "Inner loop" sigaafs a r e introduced to the main rotor flight control system a t the linear actnators which a r e an integral part of the

NOVEMBER 21, 1978 Revised MARCH 2 4 1981

P a r t 2 Section I Description FUGHT MANUAL

flight control system. Wuter loup" signals, to reposition the cyclic stick, are intrcduced to the main rotor flight conbol system by cyclic stick trim actuators. '!Inner loop" sim are introdxced to the tail rotor flight control system a t the tail rotor actuator. "Outer loop" sigcrals, to reposition the tail rotor pedals, a r e inbodaced by a yaw damper trim actuator. With an AFCS Phase IU equipped aircraft, when the cyclic stick .trim is ON and the yaw trim is OFT, the AFCS is functionally identical to a Phase 11

systern.

Pitch and roll signals a r e provided to the pitch and roll channels by two attitude gyros. Abwe 50 KIAS an airspeed transducer also provides a s i g d to the pitch channel. Tke pilot may select a desired trim attitude reference with either the cyclic TRIM REL button or the AFCS BEEPER m.The inner loop of the AFCS attempts to correct for any deviations from the reference attitude, however, if additional control authority is required, a signal to the cyclic stick trim ac-tors (oakr loop) repositions the cyclic stick.

YAW CfIANNEGHEADING H O W Beading hold is propided a t all airspeed with feet off the pedals. Signals to the yaw AFCS are provided by rrPo paw rate gyros aad a directiondl gyro. When the helicopter deviates from the selected heading, c o ~ t i o is n initially provided by Ute inner loop paw AFCS actuators a t the tail rotor semo. Additional authority beyond inner loop capability is provided by outer loop signals ta the pedal clamper trim aclnator to reposition the pedals. YAW CHANNEL

- TURN COORDmTION

Turn coordination is aoaifable abwe 60 gZAS with feet on the pedals. Signals to the yaw AFCS are provided bp two yaw rate gyros, trPo lateral accelerometers, and the roll AFCS. Inner loop and o u k r loops function together as described in HEADING HOLD. AF(S CONTROL P A .

See Figure 1 - 2 3 k AFCS INDICATOR See Figure 1 - 2 3 k See F'igares 1-16, 1-18, and I-23A The cyclic stick trim switch on the master switch bas three marked p i t i o m : AUTO, ON, and OFF, AUTO =t be used f o r Phase III to p m vide automatic cyclic stick positioning. ON will provide Phase If operation only without anbmatic cyclic stick positioning.

NOVEMBER 21, 1978

Revised JUNE 28, 1984

Part 2 Section I Description FLIGHT MANUAL



0000a0000 000000000 OOOOOUOOO 000000000 000000000 000000000 000 000000

No. 1 COMROL

NO. 2 CONTROL

S66626clm

F'igrae 1 4 3 k AFCS Pbase III Control and Circuit Breaker s EfOVEblBER 21, 1978 Revised Suly 26, 1982 1-44B

Part 2 Section I Description FUGHT MANUAL YAW TRIM SWITCH

The yaw trim switeh on the master switch pawl has bvo marked positions: ON and OFF. ON mast be ased for Phase III to provide automatic tail r o t a peddl positioning. OFF renders outer loop pedal motion and Phase III yaw functions inoperative,

I

AUTO m - D N CAUTION LIGRT

This light will go on when a malfunction has occurred in any of the channels. To determine the m a l f a n ~ t channel, i~ refer to Figure 3-1-0 in Part 1. BasicaUy, the procedtae is to place the yaw trim OFF. If the light stays from AUTO to OX. ON, turn the yaw back OK am?place the cyclic AFCS A/S ADVISORY LIGRT This light w i l l go on to indicate that the airspeed portion of Phase III AFCS is opexational. The light should be on a t airspeeds greater than 45 to 50 KIAS and out a t lower airspeeds. The light will, hawwer, remain on a t lower airspeeds if the slowdown is done a m n s t longitudinal trim force.

A test switch is provided to checkout the AFCS Phase IU "AUTO SEI7JT-DN" frmction. This frmction is enabled only while the a i r c a f t is on the ground the cyclic and is interlocked through the weight-on-wheels switch trim master switch in "AUTO" position, moving the test switch canses the cyclic stick to move in the selected direction. The motion w i l l be stopped after approximately one second and simultaneously the "AUTO SHUT-Dhn' light shou..d light

F'ull AFCS Pbase IKI capabilitp is a-le with AFCS Phase II on and aperating, cyclic stick trim switch at AUTO, and the yaw trim switch Oh'. At hover and in slow flight 0 to 45 KIAS, this provides pitch and roll attilnde retention and heading hold with feet off the pedal switches. Above 45 to 50 KlAS an airspeed hold function is introdrrced, and a t airspeeds above 60 IIIAS, tnrn coordination becomes a-e. AttitPde retention has the capability of maintaining the trimmed pitch and

roll attitade throu@mut the certified envelope. To accomplish this, especially during maneuvers where substantial trim changes would be expected, small cyclic stick displacements may be observed. This is normal and chaacteristic of an auter loop system. As in AFCS Phase II, the pilot can maneuver the aircraft by flying against trim o r selecting a different attitude reference using the trim release or cyclic beeper sysDem. Unlike some AFCS Phase II configured airmaft, the Phase IIt cyclic beeper -stem contirmauy cbanges the attitade reference without the 3 second time limit as long as it is actuated. This Phase III configured aircraft beeper function is inhibited if the inner loop command exceeds one half of its full authoritg limit

NOVEMBER21, 1978 Revised July 26, 1982

I

Part 2 Section I Describtion FUGHl MANUAL

The airspeed hold function is a contributing factor to pitch attitude retention and is on line a t airspeeds above 45 to 50 KViS when either of the following conditions a r e satisfied: The pilot moves the cyclic stick longitudinally against trim. or 25 seconds have elapsed since 1ongitPdiaal beeper trim o r the cyclic stick lrim release as last actuated. The airspeed function is insystem without compromising the tended to entrance the AFCS Phase pilot's ability to fly the aircraft normally using attitude. It has the capability of holding airspeed about a trimmed point whether level. climb, o r descent and then once eslablished wiU mainmin that airspeed throughout changes in flight path. The longitudinal stick force method of engaging airspeed hold is best used when established a t a trimmed condition Conversely, attempting to s e t an airspeed reference when accelerating/decelemting past the desired airspeed is not recommended. Tbrn coordination, available above 60 KIAS with feet resting on the pedal allows the pilot to make heading changes. automatically coordinated, using only roll cyclic. Small heading changes snch as those required drrring a precision izrstrument a p p r o a c h are easily accomplished using equally small roll angles. The pilot must allow the pedals to move in outer loop while keeping the peda3 switches depressed, and a y apparent pedal force while turning is indicative that the pedals a r e being restrained. switches,

-

Beading hold, available at any airspeed, can significantly reduce pilot workresults, heading bold engagement should be accomplished load. For only when tke aifiraft is trimmed at the desiheading before Wang feet off the pedal switches. The presence of a turn upon engagement will z s d t in a miscoordbated trim that is easily corrected with a small amount of IOU,and the use of roil beeper trim is recommended. The heading hold reference wi33 not change if this occurs.

'Lhe tricycle landing gear ccmsists of a full swiveling nosewheel assembly and -0 main fanding gear assemblies equipped Faith hgrdaplic b-s. Hydratrlic pressure from the second stage servo hydranlic pmnp is used to extend o r retract the landing gear. The extension-retaction is cmtroUed electrically an power from the dc essential bas through a circuit breaker

NOVEMBER 21, Revised MAR-

1978 2 4 1981

'

Part 2 Section I

Description

Figure 1-24.

Landing Gear Control

marked 19G GR COhT. Indicator &his are connected to the dc essential GR POS LTS. An emergency extension bus by a circuit breaker marked system uses compressed a i r to extend the gear if the normal System malfunctions. All fanding gear controls and position indicator lights are contained on a marked L9G GEAR ( T i 1-24)on the right side of the iastnnnent .

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The landing gear control handle marked NORMAL UP and DOWN, tbree green indicator lights marked L (left), N (nose), and R (right), and a red indicator light marked UNLKD, are on the landing gear costrol . The control handle m a be pulled up to move it 'to the UP or DOWN position. The indicator light will be an whenever the gear is in the retractian red o r extension cycle, o r whenever one o r more landing gear assemblies is not locked in the up o r d m position. Each green indicator light will be on only when the corresponding gear assemhly is down and 10-d. W four lights sbodd light when the cxmtim-advisory TEST Spritch is depressed.

LANDING GEAR UP WARNLNG LIGHT

-

PRESS TO RESET warning light capsrrle is installed an the instrrrment . The light wRl go on and an intermittent 250 Ek tone witl be heard in the headset if the landing gear is up zud airspeed is less than about 40 o r 60 KIAS depending on the parti& airspeed switch instaUed. The tone may be eliminated by pressing in the warning light capsale but the light will remain on until either the hading gear i s lowered o r airspeed is increased to w e r 40 or 60 KMS. A second v a m b g light capsule may be on the copilot's side of the hstmment . A LDG GEAR UP

NOVEMBER 21, 1978 Revised MARCH 24, 1983

I 1

Part2SecmnI Description

RIGHT MANUAL

-

EMERGENCY LANDING GEAR SYSTEM A pneuma?ic system can be used to lower the landing gear if the hydrauEc system malfunctions. The cantents of an air container charged to 3000 psi air pressure is discharged into the system to pressurize the down side of the hydanlic pistans. The air charge also repositions a Pafve which byes the main landing gear oalve and apens the retarn flow line. The system is operatgd by turnkg and then pulling .the red T-handle marked

RMERDN-?'IIRN&mfLLanthelandiPggearcarrtrolpaael. Afterthe system is used, the T-handle and the valve which byes the main &zdhg gear valve (right wheel well) must be manmlly reset, and the air co&iner recharged. The air container in the electronics camparbent has a cfiarging valve and pressare gage.

S e mpitot-static spstems for pilot's and copilot's instnrments provide pitot and W c pressure to the airspeed indicators. Tmasdncers a . switches ia the pitat-static system are nsed as sensors for the landing gear warning system, pitch bias system, and AFCS systems. Electrical heatgrs preventtkeformati~~~oficeontheheads.

TAZL CONE STATIC

STANDBY COM

Afree-air-indicator onthe -sideofthe meni below the windshieid Micaies is dire* OAT, ON-OFF light

in

m-

pilot's Celsius. An

above the standby can.

FUGHT MANUAL

Part2 Section1 Description

Attitilde Indicators

Mitade indicators are provided as optional equipment for the piiot and copilot. A second standby indicator is also installed on the pilot's side of the . AU indicators are idenfical and contain a pall-to-cage Bnob, a ahplane, and an OF3? flag. The pilot's height adjustment for the and the standby indicators are connected to t k dc essential bus by circuit breakers marked A?T PLT and ATT STDBY. The copilot's indicator is connected to the No. 1 dc primary bus by a circuit breaker marked ATT LT.

Directional gyros are provided as optional epipment for the pilot and co- . pilot. Both indicators are identical and contain a push-to-cage and heading set knob and an off flag. The pilot's indicator is connected to the dc essentid brrs by a circuit breaker marked PLT HDG. The copilot's indicator is connected to the KO. 1dc primary bus by a circuit breaker marked LT 3DG.

Tran and Slip Indicators

Tmm and s l i p indicators are provided as a p t i d eqrripment for the pilot and copilot. The pilot's indicator is connected to the dc S e n t b l bus by a circait and the copilot's indicaior is mnneckd to the breaker marked T & S No. 1 & primarg bus by a circait breaker marked T & S LT.

The C-14A com system combines the f-s of a flux valve and a directid~toprovidesbbilizedmqpefichesLdirtginf0nn;Ltion.Thesystem may also be used as a fnx directional gpzo at the option of the pilot. Magnetic heading information is presented on the rotatiag dial of the d o magneiic indicators. Electrical power is supplied to the system from the No. 126 pac primarp brrs tlmmgh a circuit imaker m v k e d COXP. ContmlsconsistirrgofaFEZEE~VEDaPdahLANSYNCswikh,andan

aanrmciator, are onasnaallpactelmarkedCOMantheaftendoftbe' center console. To opeate as a gyro stabilized magnelic com, the FREX-SLAVE switch is placed to SLAVE. A symhan&r automatidy aligns the directional gyro to the magnetic heading at a fast slave xi&. Proper aLignment is iudicated when the annanic&or meter is centered, The MAN SPNCmayalsobeusedtosynchronizethegyratothemagnetkheadingby moving the switch to the same position (+) or ( ) as in&by the anmmciator pointer tmtil the pointeris centered. To ope* as a directionalgpro,theFELEE-SLAVEDswitchisplaeedtoFREE. -MAN SYNCswitthispfacedtotheleft(+)totoe, ortotheright(-)to&m e the com heading indicated on the radio magtetk indicator. Tke comspstem iscannectedtotheNo. larixnarybrrsandthf!No. 1 26 vac bPs by circnit breakers marked COMP.

NOVEMBER 21, 1978

Part2SectionI

Description

The face of the aUitude indicator consists of a stationary minilta# airpbne tbe helicopter, a bank angle scale, bank index, and a moving r e -0-colored sphere with a distinct horizontaZ line d i v i b g the light color above the darker color below. A warning flag mzutbd Om will appear on the tace of the instrament when the indicator is inoperative, and for aborrt 68 seconds after ac power bas been applied to the spstem. A pitch trim h o b onthe right ofthe case may be usedto deflectthe horizontalline afthe sphere upward o r downward to trim for ihe pitch attitude of the helicopter. A roli trim knob an the left of the case mag be ased to tilt the horizontal line to trim for the roll attitude of the helicopter. A turn and slip indicator mouxxted below the attitude indicator indicates me-of-turn and slip o r &id rate. The 'pilot's indicator opemes from the No. 2 ac primary bns the Xo. 2 26 bns and the copilot's indicator operates from the No. 1ac primary bus and the No. 126 vac bcs thrcmgh cjrcnit breakers marked A'IT PLT and A'IT LT.

-

The course deviation indicator is.with VOR equipment to indicate deviation from a selected omni ange =dial o r X I S localizer. The OBS h o b is used to rotate the com card to select the desired omni radial. The v e r h d coarse deviation pointer indicates the deviafion fram the selected omni ange radial o r the course signal frrrm the XIS bcalizer. The glideslope if installed, to indicate vertical deviatioa from an on c o ~ r s glideslope e signal, A TO-FROM window indicates the heading of the helicopter (to or from) the selected station. LOC and GS warning £lags m e of m e signals.

The a d i o magnetic indicator presents heading infomaatha on a mtating com . - card which is coupled tothe C-14. com system, The card, wabmmg an azimuth scale, is mtated by the com system to indicate themagneticheading~ftheheLiwpter~afixedindexatthetopof the h&cator. A single-bar and a doable-bar pointer provide magnetic bearing informitb to VOR or ADF statians provided by W v e r VOR or ADF eqrripment may be insblkd. The radio magnetic indicaior i$ connected to the dc essential b.pS and the No, 126varimarpb by a circuit breakersmarkedRMX.

I

The wheel brake sgstem is a self-canfair& hydradic system that oper;rtes oa~Pliressraefromtoe-~bagecylindersonthepitot'stail rotorpeda3.s. Aparldngbrakehandleisonthe~sideoftheaftcantroL nsole. To set the parging brake, depress pedals, pull x p on brake handle, then release pedals. To release bake, depress pedals and the handle will snap down. Brake pedals may aka b installed on the copilot's tail mtor pedals on some ~ c a p t e r s .

NOVEMBER 21, 1978 Revised MARCH 16, 1979

I

CAUTION/ADVISORY

LDO 0EAR WARNINQ LIQHT

PIIE9S~fOtIEOFF

PflESS.lOllE OFF

MASTER WARNINQ

PRBSS lOllE OFF --

-

-

MASTER WARNIN0 ON HELICOPTERS MODIFIED BY P N 7 0 0 7 0 - 5 5 0 4 9 - 0 1 1

* CAPSULES FOR

OPTIONAL EQUIPMENT

onr

II I

Part 2 Section I Description

~ NFLIGHT K @MANUAL RH~~Y~&

CAUTION-ADVISORY The caution-advisory pmeS on the instrument contains five rows of light capsules. The red warning lights or amber caution lights indicate certain system failures or unsafe conditions, and the green advisory lights call attention to certain non-critical conditions. Each light capsule is controlled by individual circuitry through circuit breaker (s) on tne dc essential and/or dc primary bus(es) . When a light goes on, it will remain on as long as the condition exists which caused it to go on. When the TEST button on the is pressed, all lights should light on the cautionadvisory , master warning s, landing gear , fire extinguisher , digital Tg repeaters and if installed the NO GO lights on the AFCS control s, and the red WARN light on the flotation gear . Power for the test function is from the dc essential bus through a circuit breaker marked CAU ADVSY. A BRT, DIM switch on the pznel is used to control light brilliance.

I

I

MASTER WARNING An amber master caution light and three (four on helicopter modified by PN 76070-55049-011) red warning lights, to indicate engine fire or failure, are-on this (Figure 1-25). To direct the pilot's attention to the caution-advisory when a warning o r caution light goes on, the MASTER CAUTION-PRESS TO RESET light will also go on. After the condition has been noted, the master caution should be reset to &low it to light again if another caution light shouid go on. The master caution light does not light when an advisory light goes on on the caution-advisory . For a description of the FIRE, and the NO. 1 and NO. 2 ENG OUT warning lights, refer to paragraph titled EMERGENCY EQUIPMENT. LIGHTING SYSTEM EXTERIOR LIGHTS Positior- Lights The three position lights are controlled by an EXTERIOR LTS - POSITION switch on the master switch (Figure 1-9). The switch has marked positions ON and OFF and is connected to the No. 2 dc primary bus by a circuit breaker NOVEMBER 21, 1978 Revised July 10, 1987

SA 4047-76-1

A ~ ! ~ ~ M wPart ~& 2 Section I Description FLIGHT MANUAL

marked LIGHTS POS. The red (left) and green (right) position lights are on the stabilizer tips. The white position light is on the aft end of the tail rotor pylon. Anti-coUision Light The anti-collision light, mounted on the top of the tail rotor pylon, is controlled by an EXTERIOR LTS - ANTI COLL The switch switch on the master switch (Figure 1-91. has positions marked ON and OFF. The light circuit is connected to the dc essential bus by a circuit breaker marked ANTI-COLL LT.

NOVEMBER 21, 1978 Revised July 10, 1987

SA 4047-76-1

Part2Secti0nI

F U M MANUAL

Description

Controllable SearWM A full s w i v e k g , tiltable seas~bLi@ is installed on the bottam of the f s e l a g e forward of the nosewheel. The searchlight may be tilted doffaward and turned in any direction by electric motors c o m e d by switches in the pilat's compartment. The searchlight circuits are connected to the dc essential b by circuit breakers marked. SEARCHLlGEIT PWR and COhT.

-

Searchlight Master Switch. The SRCH LT master switch on the master switch (Figure 1-9) has positions m e e d ON, OFF, STOW. ON toms on the light and energizes control switches on the cycLic stick grips. STOW retracts the light to the stowed position and tarns it off. OFF trvns off the light but does not retract it. Searchli@ Conbol Switch. Thmnhmitches on the cyclic control stick grips marked SLT T m have four positions ma,-ked FWD, AFT, L, and R, with directional arrows for each. When the searchlight master switch is ON, the conttol switches are moved to any of the four positions to direct the searrhligM beam to the desired direction. When the desired direction is obtained, the switch is released and i t will return to the center position.

A fixed faading light n w be M e d in the nose of the helicopter in lieu of the contmllable spotlight. The fixed light is conhdled by the same switch Uid is used f o r the spotlight and is connected to the dc essential bas by the circuit breaker marked LDG LT PWR.

A rotary -3'T

INSF LTS, PLT switch on the center cansole tmns an and c ~ t h e i n t e n s i t y o f t h e ~ h s t m m e n t l i P M f ?.helightcircuitis i. connected to the No. 2 dc primary bus by a circuit breaker marked LTG DIM-PLT INST. When capilat's instrrrments are installed they are controlled by a similar switch marked LT located aft of the pilot's switch. The copilot's irrstrnm- light c-t is connected to the No. 1 dc primary bus by a circuit breaker marked LT INST DIM.

-

A rat;up NON-FLT lIWT light switch on the master switch (Figure 1-91 tmu an and contrafs the h b n s i t y of the nan-fright 'lipfrts. The light circuit is cop~lecf~d to the dc essential bus by a circuit breaker

-

marked NON-FLT INST LTS DIM.

NOVEMBER 21, 1978

Part 2 Section I Description

~IMBS/MT~& FUM MANUAL

Lower Console Lights A rotary LOWER CONSOLE light switch on the master switch (Figure 1-9) turns on and controls the intensity of the lights on the lower console. The light circuit is connnected to the No. 2 dc primary bus by a circuit breaker marked LTG DIM-CSL.

Pilot's Compartment Secondary Lights Six lights recessed in the instrument glare shield provide a secondary source of lighting for the instrument and control console. The lights are controlled by a GLARE SHIELD switch on the master switch (Figure 1-9) with positions marked BRIGHT, OFF, and DIM. The light circuit is connected to the dc essential bus by a circuit breaker marked LIGHTS-SEC. Standby Com and OAT Gage Lights

I

I

Lights for the standby com and OAT gage are controlled by individual light switches mounted on the bracket directly above the standby com. The switches, marked CMPS and OAT, . have positions marked ON and OFF. The com light is connected to the dc essential bus by the LIGHTS-SEC circuit breaker. The OAT light is connected to the #2 dc primary bus by the LTG DIM FLT INST circmt breaker. Cockpit Floodlight A cockpit floodlight, a flood light control switch, and a rheostat are on a (Figure 1-26) above the circuit breaker s in the aft center section of the cockpit canopy. The light is controlled by the switch marked CONT, OFF, and BRT. Light intensity may be adjusted by tPrning the rheostat when the switch is at CONT. The floodlight circuit is connected to the No. 2 dc primary bus by a circuit breaker marked LIGHTS-FLOOD.

Cockpit Utility Light A cockpit utility light and extension cord is mounted on the side of the cockpit floodlight . The light is turned on and off by pressing in on the back end of the light casing. Light intensity may be varied by tnrning the back of the casing. The light circtlit is connected to the dc essential bus by a circmt breaker marksd LIGHTS-UTIL,

Cabin Dome Light A cabin dome light,& mounted in the ceiling of the cabin. The Eght is controlled by a switch on the master switch (Fignre 1-9) marked CABIN. The switch has positions marked BRT , OFF, and DIM. The light circuit is connected to the No. 2 dc primary bus by a circuit breaker marked CABIN. NOVEMBER 21, 1978 Revised July 10, 1987

Part 2 Section I FUGHT MANUAL

Description

S 53369 (8)

Figure 1-26. Cockpit Flood Light and Utility Light

NOVEMBER 21, 1978

Part 2. Section I Description

FUGHT MANUAL

Baggage Compartment Light A light in the baggage compartment is switched on whenever either baggage compartment door is opened. The light circuit is connected to the dc essential bus by a circuit breaker marked LIGHTS-BAG .

EMERGENCY EOUIPMENT AND SYSTEMS

Infrared flame detectors in the forward and aft sections of each engine compartment transmit a signal to an ampMier unit if there is an engine compartment fire. The

NOVEMBER 21, 1978 Revised May 15, 1992

FLIGHT MANUAL

Part 2, Section I Description

amplifier unit lights the FIRE-PRESS TO RESET TONE (#1 and #2 ENG FIFE PRESS TONE OFF on helicopters modified by PN 76070-55049-01 1) warning light on the master warning s, m s m i t s a continuous tone signal to the pilot's and copilot's headset, and lights the affected engine quadrant T-handle warning light marked NO. 1 FIRE or NO. 2 FIRE. The tone signal can be reset for other emergencies by pressing in on a master fire warning light capsule. The waming lights and T-handle light will remain on until the fire is out. The No. 1 and No. 2 engine fire detector circuits are connected to the dc essential bus by separate circuit breakers marked FIRE DETECTOR - NO. 1 and NO. 2. A test circuit is provided to test the system for proper operation; however, it does not test the infrared sensors. Fire Detector Test Switch A switch marked DET on the overhead control (Figure 1-27), has positions marked AFTIBAG, NORM, and FWD. When the switch is moved to AFTIBAG or FWD. both engine fue waming lights on the master fire warning s and both T-handle lights should light and the continuous tone signal should be heard in each headset. Pressing the fire warning light capsules on the master warning s will reset the tone s i p l s for other emergencies. Placing the switch to NORM will switch off the lights and ann the system to indicate an actual ensine fire. The test circuit is connected to the dc essential bus by a circuit breaker marked FIRE DETECTORTEST. The SMOKE DET BAGGAGE warning light on the caution/advisory will also go on when the switch is moved to AFTIBAG. ENGINE FIRE EXTINGUISHER SYSJTM I

A CF3Br fire extinpisher system is installed in the engine compartment.

The fire extinguisher agent is stored in two pressurized containers that are in a compartment aft of the engines. Each container has a main and a reserve dischaqe valve which will release the agent into the compartment of

0

SHORT

0

NORM

EST

AFT, BAG

M

"lME

PRIME

FWD

F i g r e 1-27. Overhead and Fire Extinguisher Control s

NOVEMBER 2 1, 1978 Revised May 15, 1992

Part 2 Section I Description

-

FUGHT MANUAL

the selected engine. The main dischazge valve of each container releases the agent into its respective engine compartment and the reserve valve of each container releases the agent into the opposite engine compartment. The engine compartment to receive the agent is selected by pulling back the appropriate T-handle on the engine control quadrant. Each fire extinguisher agent has a redundant electrical power supply. The No. 1en,gine system is connected to the dc essential bus and the No. 1dc primary bus by circuit breakers marked FIFtE EXT and the No. 2 engine system is connected to the dc essential bus and the Fio. 2 dc primary bus by circuit breakers marked FIRE EXT. Engine T-Handles The engge control quadrant (Figure 1-4) has a T-handle for each engge, which is normally in the f u l l forward position. If an engine compartment fire i s indicated by the continuous tone signal in the headsets, and the of the master warning light and the T-handle light of the affected ligengine, the lever is pulled full aft. When the T-handle is moved full aft, the engine lever and the fuel lever are moved to OFF, the dc generator is switched off, and the fire extingcruisher is armed to fire into the affected engine compartment. Fire Elxtin,gher Switch

I

The FlRE EXT switch, on the windshield center column above the standby com, is used to release the fire extingcruisher agent from one o r both containers into the compartment of the en-e selected by the T-handle. The switch has positions marked RESERVE and OFF. After selecting the e n m e compartment with the TThandle, the MAIN position will release the agent through the main valve of the corresponding container. The RESERVE position releases the agent through the reserve valve of the opposite container. Fire -pisher

Test Switch and Indicator Lights

The EXT test switch, a red WARN indicator li*, and a green TEST light The lever lock type switch has three positions marked SHORT, OPEN, and NORM. The switch must be pulled out to move it out of the NORM position. To test for an open circuit in the system, place test switch to OPEN; the red WARN l i t should go on. Pull both T-handles back about 1 inch o r more and place the FIRE EXT switch to MAIN then RESERVE. The green TEST light should go on in each position if there is no open circuit. To test for a short circuit in the system, place test switch to SHORT with' both T-handles pulled back. The red WARN and s e e . TEST lights should go on, the preen light should stay on when the FIRE EXT switch is moved to MAIN, then RESERVE. Normal operation is restored by moving the T-handles full forward and returning the test switch to NORM. Both the red and green lights should then go off. The indicator lights should go on when the caution-advisory TEST switch is pressed.

1 a r e on the overhead control .

NOVEMBER21, 1978 Revised OCTOBER 14, 1980

FLIGHT MANUAL

ENGINE FAILURE W-G

Part 2 Section I Description -

sXsTEM

TheNO. IandNO. 2 ENGOUT -RESfiTONEwamingligfitsonthexnaster warning (Figure 1-25) will go on and an alternating tane (550 and 700 %) will be heard in the headset when N1 for the correagine drops to 59%. The tcme signal is reset for ather emergency cnnditiarrs ky pressing in on the warming light capsule. The alternatirrg tane signal is -ve when the helicopter is on the grotmd; however, the warning lights

fIZ

NOVEMBER 21, 1978 Revised M y 26, 1982

Part 2 Section I Description NGHT MANUAL

will be cm whenever N1 is less than 59%. An engine-out ground test switeh on the right side of the c a t e r console is used to test the tane signat when m the ground. - The alternating tone will be heard when the switch is held in to TEST. Pressing the el and 32 ENG OUT warning capsules will shut off the audible w a ? . The w a m b g circuit is connected to the dc essential bus by a circuit breaker marked A m Wm.

A SMOKE DET BAGGAGE warning light an the caution-advisory and the master caution light will go an when smoke is sensed by a detector an the forward bulkhead of the baggage compartment. The smoke detector circuit is coxmected to the No. 1dc primary bus by a circuit breaker marked SMK DET. A test circuit will Light both the warning light and the caution light when the fire detector switch is moved to AFT/BAG.

Tivo hand-operated fire tz&i~@~&s are provided. One is secured to the right side of the flight controls enclosure in the pilot's coqartment and the other in the cabin. FmsTAmKrI' A first aid kit is along the left side of the fiight control housing.

E m e x w y lghting is provided by a power pack with a two-bulb light mountedmthe&ince.i&g. ItisccnrtrdlledbyammarkedEMERLTSon positions ON, ARM, and OFF. On the master switch p a n 4 with some installatians there are two one* l i g h t s separated and mormted away from the power pacL At ARM the iight (lights) w i U go an when no puwer is snpplied to the No. 1DC PRlMARY BUS. The system is protected L X , an the NO. 1 DC PRIMARY BUS. by a circuit breaker marked If the switch is left at the light wiLl automatically go on when power is removed from the dc primary bus, To prevent this,place switch OkF before & & b g dawnlast engine or DC GEN.

AnemergencylOCatOrtansmittet~~installedasoptianaleqaipment, It is attached to tbe W t gide of the fLight cmfrols enclosme in the pilot's conpatmerit. The ELT may be used to h-ansmit a gignal an the intern&-a1 dissesshxpmciesof 121.5 and243 MHz. Powerto transmit is froma self - I-. Tttesignalcanberecei~edatadisbnceofabout 100 milesatasearchaltit&eoflO,OOOfeet. Aswitchanthetansmitter has three positions marked ON, OFF, and ARM. ON energizes the traps'

mitter. B R M a a t o ~ y e n e r g i z e s t h e ~ w k n a S G i x c p c t

.

load is encountered. OFF dewenergizes the transmitter, If the transmitLer isactmatedbythe5Gforce(MZM), ~ s w i t c h m P s t b e ~ e d O N , t h e n O F F , to de-energizethetansmftter. The is portable and is easily ren;aved from the helicopter, Discormect the antema cable and release the strap atk&ing it to the flight cantrols enclosure. A£ter removal, capnect theportableantema, s t o r e d o n t h e s i d e o f t k t -'ter case, tothetopof the transmitter, The tmnsmitter is then energized by placing the switchON.

NOVEMBER 21, 1978 Revised 3uly 26, 1982

Part2SectionI Description SOMC IDcA'mR An underwater sonic locator system may be installed as optional equipment in the forward cabin overhead, left side of the Micopter. It automatically radiates a pulsed signal to aid in locating the helicopter after a crash landing m fresh or salt water. Power to transmit the signal isfrom a selfcontained battery. The operating depth of the signdl is from sea level to 20,000 feet, and the operating life is 30 days. The signal can be detected by portable receivers eqnipped with directional hydrophones.

FIRTATION SYSTEM

For emergency water landings an emergency flotation system (Figure 1-28) may be installed. It is recommended that during an approach or takeoff over water that the pilot place the floats switch to ARMEI) when the airspeed is less than 75 KLAS. To prevent inadvertent actnation of the flotation system, place floats switch OFF daring ground operation. It is designed to provide flotation time for the immediate evaation of personnel and survivaleqaip ment after ditching the helicopter. The system is actuated electrically by the pilot jast before ing the water or at speeds up to 75 gIAS. A notation bag is installed in each main landing gear wheel well and two are installed in cornparbents next to the nosewheel well. Each bag is divided into two comparbxnts. The contents of two air bottles are used to innatP the nosewheel flotation bags with helium or nitrogen. One bottle discharges into the forward compartment of each nosewheel flotation bag and one into the aft compartment of each to provide balanced flotation. Two additional botUes a r e ased to ixdhte the main landing gear float-bags m tke same manner. Float irdlatian time is witbin 5 seconds when helimn is ased as the lmfktian medium. Nitrogen is an approved sabstitute for helium but requires ? to 10 secands for inflation. Explosive liPBs are disckuged to fully open the main landhg gear doors when the system is activated. The floats shodld be discharged with laading gear extended, Redmrdant electricdl circuits for mFlatirm are connected to the dc essential and No. 2 primary buses by three circuit breakers on each marked -AT.

I

rnfhtian switch is pressed. After completing test, Aeatw iPfktron pkce arming switch OFF and test switch tD NORM. EEATlNG AND Vh"L?LATING SY3TEM

The 25,000 BTU heatirrg system n & ~ shot air fmm fize engine compressor sectionswith cold air dxamn into a mixing valve by ventmi action. The air then es through a sotmd suppressor and into the pilot's conparbent and cabin heater drrets. Some ducted air flows acmss tire windshield for defogging. The heater can be operated only when b t h engines are operating. The heater will automatically shut o£€if one engine is shat &mn, if bleedair pressure is lost, o r if bleed-air pressure o r temperature is tDo The ventihling system uses the heater ducts and an electrically-aperated blower to circrrkte oatside air h r r g h o u t the fnterior of the helicopter. confipl cizmits are connected to the No. 1dc primarg Eeater and bus by circcrit breakers marked HEAT CONT and VENT COhT. A marked HEATER CONTROL ia the copilot overhead c o n e all controls fir the operation of these systems. An additional venh'htinp system per P/N76070-20045 cansists of a I1.lanaally consoliable inlet in each door and two exhaPst oalves, one on each side of the helicopter near the aft cabin bulkhead. This ventifating m r n &d\d be bckd cbsed for proper heater system operation. To do this, dose all door vents and press the lock arm on boa exhast valves, one on each side of the Micopter in the aft end.

-.

HEATER CONTROL

I

The heater contm1 (Figure 1-28) axtabs a bleed-air switch, a heater on-off switeh, a temperatare control switch and a blower switch. The BLEZD AIR switch bas trPo pasitions marked ON and OFF. ON opens a bleed-air valve wbkh directs engine compressor air iub the heater spstem. The bleed-air valves w i l l open oply when both eugines are opemting. OFF cbses the bleed-air valves. The BEATER switch has posr+tnnP mazked ON and OFF. ON permits airtlow h r g h the mixing valve and into the interior of the heater. OFF closes be valve. when tfig heater switch is ON, cabin temperatrrremapberegWkdbytraaingtheTEBdPCONTsFpttChfDthedegfred setking. Thenblower is eontmlled by the BLOWER switch When the hoist is ope-, the blower, ifon, will be shut off by a relay interlock c m t .

The ECU provides &matic amtm1 for beating or cooling of the interior of the helicopter b mainlab a consbut select& temperatare. The system uses. bleed-air from the engine compressors for hprZtillp and for the aperation ofanaircondithingunit. Thesy&mcanbeopexatedodywhenbothenginesareoperating. S e r r s o r s 9 C r i U a a z b ~ ~ o f f t b e s p s t e m i f o n e engine is shut down, if bleed-* pressure is lost, if bleed-air pressme or is PDo high, or cabin inlet duct t e m p e m is tDo high. Cootrols for operation of be system azle on an ECU CONTROL in the Overbad. PuweraadEontro1cirmitsforoperatiosofthe~are cmnettedto theNo. 1drimarpbusbyEircrrftbreakersmaxkedECU

-

PWRaadECUCONT. WhentheboWisoperatingarrdtheECUis.on,a relay c m t will s h k off the blowes. The P/N 76070-20045 - interlock venbk&g system s b d d be locked closed for proper Em operatio4. To do NOVEMBER 21, 1978 Revised Jamrarg 17, 1983

1-56A

I

this, close aU door vents and press the lock arm on both exhanst valves, one an each side of the helicopter in the aft end.

The ECU control (Flgure 1-29) contains a COhT SEL switch with psitions MAN AUTO OFF, and two TEMP CONT switches, one for xaannal temperature control, and the other f o r automatic temperature contro1. The m d position is used only when the automatic mode malfanctiops. When the CONT SEL switch is placed to MAN, cabin air temperatPte is controlled manually by use of the MANtTAL switch marked COOL - OFF - BEAT. The switch is spring-loaded to the OFF position and governs the amount of hot or cold air by the length of time it is held at COOL or HEAT. When the CONT SEL switch is placed to AUTO, cabin temperature may be regplated by the rotary switch marked AUTO. A two-position BLOWER switch marked ON OFF is below the ECU control . This switch provides air circulation withoat use of the ECV.

-

I

-

-

ECU OPERATION

IWhen ape-

in antomatic coolmode, the ECU responds to a large step increase in selected cabin temperature by.going into full heating mode (160°F) to produce the newly selected cabin temperature quickly. This featare is typical of current Hamilton Standard control systems. It is recommended that the following operatianal guidelines be adhered t o in order to ' enger comfort during the above transient condition: '

1. In ADTO mode of operatios do not initially select full cold. A +ely 120° dockwise from full cold will selector rotation of app provide a comfortable sdetted tempe-e of appraximatdy n ° F . 2.

If a significant (5O to 10°F)increase in cabin et-

is re-

qaested while in cooling mode tarn the selector control grlly in

steps of 30° with at least 1 minute between steps.

The manual mode is to be ased only when the arrtomatic mode malfmrctians. Prolonged use of the fdu cold manual mode may canse ice to form at the inlet &/or aatlet of the turbine section of the ECU leadiug to mrit damage, A malfTmction of the ADTO mode should receive prompt maintenance action

In AUTO mode, ECU output is amhnafically adpsted to &tab and - ' - the selected cabin temper;rtrw. Inmamntmodeakitairbyvalveis movedby an open loop comB3ana signal.

NOVEMBER 21, 1978 Revised JUNE 28, 1984

Part 2 Section I Description

RlGHT MANUAL

ECU CONTROL

HEATER CONTROL HEATER

BLEED

AUTO WARM COOL

om@HEAT

I

.

I

@

BLOWER

I

Figure 1-29. Heater and ECU control R i d s

The electrically-opeatedwindshield wipers are cantrolled by a WINDSEEDWIPER switch an the master switch (Figme 1-91 with positions marked OFF SLOW FAST. The windshield wiper circuit is connected to the. dc essential bns by a cbmit breaker marked WSHLD-WIPE&

-

-

NOVEMBER 21, 1978 Revised a - Y 17, 1983

Part2SectianI Description

WINDSHIELD WASHER

The electrically-opeated windshield washer is controlled by aWINDSHlELDWASHER switch an the master switch (Figure 1-9) with positions marked ON and OFF. The 6-quart reservoir and pmnp are installed in the electronics compartment. The windshield wiper circuit is connected to the dc essenbl bus by a circuit breaker marked WSHLD-WSHR. Electrically-operated windshield heaters and glass windshields may be installed as optional eqnipment on helicopters eqPipped with an ac gene*. The pilot's and copilot's windshields are k&edby sepraie spstgms tbt are controlled by WIM)SHIELD-HEATEX switches an the master switch (Figme 1-9). One switch marked P L O T and the other marked COPILOT have positions marBed ON and OFF. Contrdllels in each system will switch the circuits on and ofE to keep temperahnx within limits. ' X k pilot's spstem wiLl always have priority aver the copilot's system. If bath systems are switched on, the pilot's system will heai fixst. When the controller switches the pilot's system aff, power will be applied to the copilot's circuit. If the copilot's system does not operate and the pilot's system operates poorly when both systems = switched on, the cause map be a malfunctioning pilot's controller. To correct this, switch off the pilot's system and the copilot's system will be restored to normal operation. The pilot's and copilot's systems aze axtrolled by cirmits connected to the No, 2 and No. 1 dc primary bases, respectively, by circuit breakers marked WSHLS HTR CONT, Electrical power circni& for each system sule connected to tbe ac manitar lms by cixmit breakers marked PLT WSHLD HTR PWR and LT Ws3LD HTB m, Windshield Hot Cawtion Li&t The WINDSHIELD HOT cauticm on the caution-advisory ( F m 1-25) iipht?; whenever the pilot's or copilot's system does not shat off the heater when the high temperatpre limit 58OC (1359F) is exceeded. When the light goes on, shat afl bath spstems. If co&inued use of w3~1dsWdheat is reqrrired, the spsterr, at fa& may be i-ed. by switching the pilot's systemanandmoaitorirrgthecautiaaligM, Iftheligixtremains&thecopilot'ssystemisatfarrtt. I € t h e ~ l i g & g o e s o n t h e p i l o t ' s s y s t e m i s atfault. COntirmeusingapetatingsystemandplacewindshieldheaterswitch of the failed system to OFF. If necessary to nse the system at fa;rzlt,the s p s t e m c a n b e t y d l e d o n a n d o f f ~ y w i t h t h e ~ m i t c hwhile , closely manitoting the d o n LigM f o r overtemp. If the WIM)SEIEID HOT cantion light remains an after the saritch of the system at fanlt is placed OFF, it is possible that the electrical ors to that system have become welded together. To eliminate the overheat condition when tbis occurs,

I

pallthethreeWSHLDH?IRPWR~breakersafthespstematfadlt. PILOTS SEATS

The pilot's and c ~ o t ' seats s have lap-iype seat belts and iwrtia-type shoulder harnesses. A vertical adjpstment lever is on the left side of the seat A *position (LOCKED--) inertia reel lock baxile is on the left side. LOCKED w i l l lock the shoulder harness in place. UNLOwill permit the reel of the ahodder harness to extend, allowing the pilot to lean forward; however, the cable reel w i l l aotomatitally lock when an impact forre of 2 to 3 Gs is encountered. Then, to release the NOVEBdBER21, 1978 Revised Jdly 26, 1982 1-59

Part 2 Section I Description

F I J M MANUAL

lock, it is necessary to move the handle to LOand then back to UNLOCgED. Storage provisions for a life vest are on the bottom of each seat:

I

DOOR OPEN CAUTION LIGHT The DOOR OPEN caution light on the caution-advisory (Figure 1-25) will go on when any of the four enger doors, either of the two baggage compartment doors o r the external power door, are not fully closed and latched. The light circuit is connected to the No. 2 dc primary bus by a circuit breaker marked DOOR WAR?% NO SMOKlI?G AND FASTEE SEAT BELT L I G m

The no smokbg and fasten seat belt lights in the cabin a r e controlled by NO SMOKING and SEAT BELT switches an the master switch (Figure 1-9). Each switch Us positions marked OK and OFF. Both light circuits a r e connected to Ute No. 2 dc primary bus by a circmt breaker marked SIGN.

CARGO HOOK A 33Wpotmd (1500 Kg) capacity cargo hook (Figure 1-30)may be installed on Ute lower fuselage strpctnre. The hook has an open thmat with a springloaded keeper to retain the load The hook may be opened to release loads electrically, or mechanically in an emergency from the pilot's compartment. A r d e a s e lever on the hook may also be used to open the hook. After re-

lease of a load, the load beam of the hook is spring-loaded to retarn to the closed position. The electrical release and advisory light circuits are wnnected to the No. I dc primarg bus by circuit breakers marked HOOK CONT and EIOOKWAEW. CARGO HOOK A#MPIG SWITCH AND ADVISORY LIGHT The CAR HOOg azming switch on the master switch (F5gme 1-9) has . positions marked ARMED and OFF. ARMED provides electrical power to switches on the cyclic stick grips which a r e used to open the hook A EiOOK ARMEL) advisory light on the carrticm-advisory fgks on when the master switch is ARMED. OFF removes electrical p e r from the hook circuit The master switch i s normally kept OFF until just before load release.

CARGO SWlT.CB AND HOOK OPEN ADVISORY LIGHT The CARGO 61Ritches an the cyclic stick grips (Figme 1-5)are pressed to

open the cargo k&. These switches are imperative d e s s the cargo hooB armingswitchisARMED. A H O O K O P E N a d v i s o r y l i g h t o n t h e ~ advisory will go on whenever the hook load beam is not locked in the dosed position.

GRom RELEASE LEVER A ground release lever an the cargo hook map be used to open the bxk.

EMERGENCY CARGO HOOK RELEASE T-HANDLE

The EMERGENCY CARGO HOOK RELEASE T-handle is onthe aft end of the center cansole. The handle is pulled upward to marmaIly open the hook.

NOVEMBER 21, 1978 Revised July 26, 1982

'

Part2SecticmI

R I M MANUAL

CYCUC STlCK GRIP

Description

I

A m 1-30. Cargo Hook NOVElVlBER 21, 1978 Revised APRIL 12, 1979

1 1-61

Part 2 Section I Description

A 600-pound capacity, electrically-operated, two-speed, atilitp hoist with 180 usable feet of cable may be installed above the aft, right cabin door. The hoist may be op-d from the cockpit by the pilot o r copilot, or by a crewmember in the cabin. When operated from the cockpit, cable speed for raising or lowering the cable is about 120 feet-per-minute. When upexat&

from a portable hoist pendant installed in the aft right side of the cabin, hoist speed m a y be selected at about 120 feet-per-minute o r 60 feet-perminute. If one dc generator should fail, cable speed is automatically reduced to abont 60 feet-per-minute. The hook assembly h a s one large and one small hook, both equipped with keepers. A I/&inch hole in the hook assembly may be used to attach guide ropes o r a safety line. A barnper assembly acts as a shock absorber when the cable is reeled in and the bumper s the bottom of the hoist. A squib cartridge cable cutter may be used in an emergency to cut the cable at the drum. A manual-type cable catter attached to a &foot chain is stowed in a pouch mounted on the cabin floor forward of the aft right cabin door. A safety harness, for the cabin hoist operator, attaches to a tiedown ring in the floor. The hoist system operates on 28 volts dc and is connected to the hoist bus in the electrical compartment by twa circuit breakers marked HOST P0R'E.R and HOIST COWTROL. The squib cartridge circuit is connected to the dc essential bus by a circuit breaker marked HOIST SQUIB. Eoist load limits and ballast requirements are in Part 1 Section I, and Part 2 Section IU of this mnrmal. If the vent o r ECU blower is an when using the hoisf a relay interlock ciritwill.~ito£f. H o l m CcJN'IROL

I

The HOIST CONTROL , on the control console, contains all switches required to operate the rrtilitp hoist o r transfer control of raising or lowering the hoist to the hoist pendant in the cabin, o r cat the hoist cable in an

emergency. Hoist Power Switch and Indieator Light The hoist PWR switch has marked positions ON and OFF. ON provides electrical power to the hoist controller as indicated by the lighiing of the green indicator light aft of the switch. OFT disconnects e l e c t r i d power from the controller. Hoist Master Switch

Thehoistmasterswitchhashvomarkedpwitions, mLQTandCREW. PILOTpermits the hoist to be raised orbweredfrom the cockpit by use of the HCUST control switch next to the master switch. CRBW transfers control of the h o i s t t o t h e h o i s t ~ inthe t cabin. Hoist Control Switch The HOIST control switch has three marked positions: UP, OFF, and DWN. Theswitch~frmction~~en~ehoist~~ibpfacedto PLLOT. The switch is spring-loaded b OFF and must be held UP o r IX)WN toaper;itgtbehoist.

Part 2 Section I Description NGHT MANUAL

Cable Shear Switch The cable shear switch with one marked position CAB= SIlEAR may be used in an emergency to cut the hoist cable at the dram. To cut the cable, lift the switch guard and place the swim to SHEAR. The shear electrical circuit operates independently of the hoist puwer switch and is connected electrically to the essential bus by a circuit breaker marked HOIST SQUIB. HOIST PENDANT

The hoist pendant is a portable hand-held unit used to control the hoist from the cabin, The pendant is connected to a coiled cable which w i l l extend to 1 2 feet and is stowed below the aft window on the right side of the fuselage. The pendant wiIl function only when the hoist master switch is placed to CREW. Controls on the pendant are as follows: Power Switch The power switch on the grip of the pendant is a "dean" type switch which must be squeezed to apply power to the hoist control switch on the pendant. When the power mPitch is released, the control switch will not function and hoist operation.is automatically stopped. Hoist Control Switch The hoist contml thumb-operated, wheel-type, switch has five marked positions: UP FAST, UP SLOtrJ, OFF, DWh'SLOW, and DWN FAST. The switch is turned to the desired position to control the speed and direction af the hoist. The switch is spring-to the OFF position. Hoist Eigh Tempemhue WarniPg Light

of t&e pendant control s h m g e compartment may be used in an emergency to cut the hoist cable at the drum. Tocutthecable, liftthe~tchgum3andplacetheswitChtoSBEAR.

A CABLE SHEAR switch in the fonvard part

A cable shear test switch and continuity test light permits ckcking the hoist squib integritp and the hoist cable shear circuitry at a redrrced current below the minimum required to fire the squib cartridge. The switch and light are mounted above the forward right cabin window. The test is

marked HOIST and has marged positions TEST and ARMED, and the test light is marked HOIST TEST LT. The circuit is tested by first placing the test switch to TEST and then placing the pilot's o r crewman% cable shear

NOVEbdBER 21, 1978 Revised OCTOBER 14, 1980

Part 2 Section I Description

Figare 1-31. UWiB Boist (Sheet I of 2) NOVEMBER 21, 1978 Revised July 26, 1982

\

Part2SectionI Description FUGHT MANUAL

H O W CONTROL

.:

CMER CONSOLE

\,.................-......./........



?:

Figrw 1-31, Utilitp Hoist (She& 2 of 2) NOVEbdBER 21, 1978 Revised July 26, 1982

I 1-65

Patt2sectianI Description

RIGHT MANUAL

switch to SBEAR. Lighting of the EOIST TEST LT indicates a satisfactory test. The HOIST TEMP warning light will also be tested (light on) when the swikh is placed to SHEAR. After testing the circ&, rethe shear switch to OFF, close the guard, then place the test switch to ARM so the shear c H t is ready for use.

To prevent accidental cutting of the cable be sure that the hoist test switch is at TEST before placing either cable shear switch to SBEAR After completing the test be sure that the shear switch guards are closed before placing the test-arm switch to ARM.

Each copy of this m a n e which is assigned to a specific aircraft (i. e. aircraft copy) must contain all data referred to herein and must be mahhined to the c m e n t aircraft coIlfigPration. Copies issued on general distribution only need not contain such da&.

NOTE

Use this section for helicopters withoot a utility hoist installed, o r for helicopters with'a utility hoist installed but carrying no hoist load. Use Section 3 I t for helicopters carqhg a load on the utility hoist, since both horizontal and lateral center of gravity must be determined.

EMPTY WEIGHT AND HORfZONTAL CENTER OF GRAVITY .LOCATION The m e n t eznpty weight, moment/lOO aad center of gram location of this aircrzft are as shawn in the Empty Weight 2nd Balance Record, Chart C herein. 1.

The current empty weight includes all items marked ( d )in the Empty Weight Cheek List, Chart A, herein. Items not included are marked (0).

2.

Chart C (Figure 2-9) mrst be kept up tq date. ff any items of equipment are added to o r removed from the aircraft, Charts A aad C mast be updated as follows:

When items are added, make the apprapriab entry in fhe Emptp Weight and Balance Record, Chart C, On Chazt A mark(v), inboth the 7nAircmY"'nd Thart C E&-y" columns apposite tke items added, dating the co1rms11s where indieabd. Add tke new entry or entries on Cbart C to the previous to&& of empty weight and moment/100 shown on Chart C and compute the new empty weight, moment/100 and center of gravity location. When items are removed fmm the aircraft, follow a similar procedure to (a) above except mark (0) in the 7n Aircraft1' column arrd(g)in ttte 'Chart C E&y" column of Chart A, and subtract the new entry o r entries from the previous totals of empty weight and moment/lW, shown on Chart C (Figure 2-9). New items to be added to the airtr;iit and not previously listed, may be '%mittenin" on the next available line in the appropriate

NOVEMBER 21, 1978 Revi~edJUNE6,I980

Part 2 Section II

d~mmmrslR FLIGHT MANUAL

Loading ~ ~ r m a t i ~ ~

aircraft compartment. Do not erase o r obliterate any Chart A entry when items are removed from the aircraft. The aircraft empty weight center of gavitg location (CG)is initially obtained by actnal wejphi(.lp (Figure2-8). The updated CG is obtained by dividing the wdated moment/100 by the apdated empty weight and multiplying the resolt by 100. The '3n Aircraft" colmnn in Chart A may be used for periodic checkbg (d the aircraft inventory. It is recommended that the first check be made prior to the first flight by a new operator.

The original empty weight and center of g m t y location values were determined by Sikorsky Aircraft, as shown in the Aircraft Actual Weight and Balance Form, Chart B (Figate 2-8), which is included ia the aircraft copy of this manual for reference. FOR AIRCRAFT WPraOm TEE UTILITY HOIST OPTTON mTALLED

NOTE Lateral center of &ravity limits permit Wiscriminsrte lateral loadingd engers or cargo if maximum fnel loading imbalance is kept below 250 Ibs wifb baggage compartment spmmetrically loabed Maximum cargo compartment and baggage floor loading is 75 pounds per square foot.

NOTE center of gr;rvity permit indiscriminate lateral loading af engers o r cargo if masmmn baggage imhlame is kept below 150 pounds with fael symmetrically loaded Allnn cargo compartment and baggage floor loading is 75 pormdls per square foot.

THE TITUTY H O R flEIJCOPTER OPERA= IFSTALLED BUT NOT OPERATED DURING =GET

m OPI'ION IN-

I

NOTE M d e d that the first enger (170 pomrds)o r the first 275 pounds of cargo is loaded to the left (port) side af the helicopter centerline, lateral center of ~ l y ~ p e n n i t ~ c r i m i n a t e ~ l o a d i n g aftheTemainderofpas~e~gersor~ifthemaximrrm fael loading imhnlance is kept bekm 50 pormds with the baggage camp;utment symmetrically loabed., The m u m PPaorm cargo compartment and baggage comp;utment floor loading is 75 pounds per square foot.

wFLIGHT MANUAL

Part 2 Section XI Lopding I~ormation

NOTE Provided that b e first enger (170 pounds) or the first 275 pounds of cargo is laaded to the left (port) side of the helicopter centerline, lateral center of gravity limits permit indiscriminate lateral loading of the remainder of engers or cargo if the maximum baggage imbalance is kept below 30 p o d with the fuel symmetrically loaded. The maximum tmiform cargo compartment and baggage compartment floor loading is 75 pounds per sapare foot. LOADING INSTRUCTIONS Obtain the current aircraft empty weight and aircraft empty weight moment/100 from Chart C (Figare 2-9) and enter each in the appropriate columns of the manifest.

Obtain the aircraft takeoff gross weight and moment/100 by adding,

to a e rrent aircraft empty weight and moment/100, the weight and momeat/100 of each useful load item to be Carried. The weight and moment/100 of useful load items is shown in Chart E Weight Moment Charts. Useful load items include fuel, oil, crew, engers, baggage and/or cargo. Do not exceed the specified compartment capcities at any time.

Determine by reference to the Center of G r a m (CG) Chart, F i i 2-1, that the takeoff gross weight and moment/lOO thus obtained is within the recommended.weight and CG limits. The aft CG limit has been adjusted to for the landing gear retraction moment/ 100 of plus (+) 33. Therefore, lading based on wheels doam conditions which fall within the limiting moments/lOO of Figure 2-1 will be within CG Limits for flight with the larrding gear retracted. '

ChecgthslttheaircaftwillremainwithinCGYimitsthro~rrtthe Wt: (a) dednct the Weight and Moment/lOC) of expendable items (e-g. feel, windshield washer fluid), (b) compute the landing gross weight and moment/lW and (c) determine by reference to the CG Chart, Figure 2-1, if the landing gross weight and moment/100 is within the recommended weight and CG limits. Bring the airmaft weight and/or halaxice within Emits if either the weight or moment/ 100 - is- outside of the recommended limits by ofbloading and/or r e dWnbutbn of useful load

NOVEMBER 21, 1978 Revised JUNE 6, 1980

5.

The CG limits are t a b w e d M and are also shown for reference

Aft

Gross Weight (po-ds)

CG Limit* (=an) 201,60 202.47 203.48 204.48 205.49 206.50 207.51 208.51 209.52 209.53 209.56 209.58 209.61 209.62 208.83 207.25 205.67 204.72 204.09

T h e aft CG limit Pafrres have been adjusted b accormt for l a d i n g gear r d r a c l i o ~ Loadings based on wheels d m conditianswhich fall wittrin the above CG limitswillbewithin CGlimitsforflipfitwiththe b d h g gear reMoment/100 for retraction of the landing gear is p b (+) 33, 6.

The maximum permissible gross weigW is 10,500 pounds.

7. Sample loading eondifions are shown in Figure 2-15.

m R 21, 1978 Revised MAY 14, 1985

1

Figtm 2-1.

Center of GraW Chart

NOVEMBER 21, 1978 Revised MAY 14, 1985

FUGHT MANUAL

-

AIRCRATT Sf* (INCHES)

NOTE

THE ~nCG. t l w n

~ n BEEN s

ADJUSTED TO FOR THE LANDING GEAR

RETRAmON HORltOKTAL MOMENT/ 100OF PLUS(*)33. THEREFORELOADINGSBASED ON WHEELS-DOWN CONDtlIONS WHICH F U WITHIN THE LIMITING C.G. OF T H S

ENVELOPE WILL BE m RETRACTED.

F -

t

LG.~ LIMITS FOR n r G m WITH THE IANDING GEAR SYWC=3)

2-2,

Weight and Center of Gavitp Envelope

NOVEB5BER 21, 1978 Revised MAY 14, 1985

REF LINE MAlN ROTOR

CENTROID MAIN UOTOR WAIN ROTOR

t

i

smn

DIMENSONS FOR WEIGHING ~ O N W H E U S O l l M S l W S FOR WEIGHING AIRCRAFT ON JACKS

NOTE

AU DIMENSIONS ARE APeROXIMATE. DIMEWSIOWS FOR WEIGHING O N W H E E L S (J AND L) M U S T BE DETERMINED BY ACTUAL YEASUREMEW.

A EUZlRONlCS COMPARTMENT

B COCKPIT

C CABIN D ENGINE AND TRWSMlSSION COMPARTWENT E TAIL F EXTERlYAL EQUtyENT

ssx7tm

F%rpe 2-3.

Diagram

NOVEMBER 21, 1978

S-76 WEIQHINQ PROCEDURE 1. THE AIRCRAFT MUST BE LEVEL WHEN WEIOHED. I T I S RECOMMENDED TllAT TtlE AIRCRAFT 8 E WEIOHED O N JACK POINTS TO PROVIDE EASE I N LEVELINO. 2. SUSPEND A PLUMB B O B ABOVE THE LEVEL PLATE (L6CAT ED ON TH E LEFT tIAND ENOER DOOR I R A M E AT STATION 176) FOR LEVELINO THE AIRCRAFT DURINO WEIOIIINO.

3. WHEN WElOHlNO O N JACK POINTS, ACTUAL MEASUREMENTS OURINO WEIOtIINO ARE NOT NECESSARY, AFTER LEVELINO THE AIRCRAFT T l l E FOLLOWIN0 DIMENSIONS MAY 8 E ADDED DIRECTLY T0 T H E WElOHlNO FORM (CHART 8):

K (HORIZONTAL DATUM TO N (HORIZONTAL OATUM TO

q q

OF FORWARD JACK POINTS) OF AFT JACK POINTS)

-

= 92.0 INCtlES 252.0 INCtlES

4. WHEN WElOHlNO O N WIIEELS, MEASURE DIMENSIONS J AND L AFTER LEVELING AND DURING WEIOHINO. USlNO TtlESE ACTUAL DIMENSIONS AND THE FORWARD JACK POINT (AIRCRAFT STATION 92.0), DETERMINE THE DISTANCE FROM THE HORIZONTAL DATUM TO TIIE q OF THE NOSE WHEEL, AND TO THE P OF THE M A l N WIIEELS. FOR CHECKINO PURPOSES T IIESE DIMENSIONS ARE APPROMMAT ELY AS FOLLOWS: IlORlZONTAL OAT U M TO TtlE IlORlZONTAL OAT U M TO TI l E

P

q

OF Tl l E NOSE WIIEEL OF THE M A l N WHEELS

'

.

= 60.0 INCtIES 257.0 INCIIES

NOTE: WHEN WEIOtlINO O N WHEELS. TllE M A l N AND NOSE OLEO'S SIIOULD B E RESERVICEO TO PROPER FLUID LEVELS W l T l l HYDRAULIC FLUID IF THE OLEO'S HAVE BEEN USED TO LEVEL THE AIRCRAFT. REFER TO INSTRUCTION PLATES FOR PROPER SERVlClNO PROCEDURES.

5, IF BLADES ARE INSTALLED, THEY MUST B E EQUALLY SPACED DURINO WEIOIIING.

LEVELING DEVICE I

1

*

FWD

(FWD

0

\~MBBOB

VIEW FJFJ ROTATED 90° CLOCKWISE

FORWARD JACK POINT

AFT JACK P01WTS LEFT H A N D SttOWN, R l O H I N A N 0 OPPOSITE

CREW, ENGER+ANDBAGGAGE CENTROIDS

AVIONICS COMPARTMENT

Standard 14Canfigmati011 ~,PasmgerandEaggageCentroids

Figrw 2-5.

NOVEMBER 21. 1978 Remised JUNE 6, 1980

--

FLIGHT MANUAL

NOVEMBER 21,1978 Revised March 2,1989

FLIGHT MANUAL

Figure 2-7.Chan A. Typical (Cont)

NOVEMBER 21,1978 Revised March 2,1989

Part 2 Section 11 Loading information

Part 2 Section I1 Loading Inionnation

A~-RB??S~~& FLIGHT MANUAL

Eigure 2-7. C h n A, Typical (Cant)

NOVEMBER 2 1,1978 Revised March 2,1989

Figure 2-7. Chan A, Typical ICont)

NOVEMBER 21,1978 Revised March 2,1989

Pan 2 Section I! Loading information

FUGHT MANUAL

Figure 2-7. C b n A. Typid (Con11

NOVEMBER 21.1978 Revised March 2,1989

SfltmmmdE FUGHT MANUAL

Figure 2-7. Chan A, Typical (Conr)

NOVEMBER 21,1978 Revised March L 1989

Pan 2 Section 11 Loading information

Pan 2 Section I I Loading Information

&!&BBSW/& FLIGHT MANUAL

Figure 2-7. Chan A. Typical (Cont

NOVEMBER 21,1978 Revised March 2,1989

SA4047-76-1

~~~~~~~ FUGHT MANUAL

Figure 2-7. Chan A. Typical t Cont

NOVEMBER 21, i 978 Revised March 2, i 989

Pan 2 Section 11 ionding Information

Pan 2 Section I1 Loading Information

SI/K@?HMI//& FLIGHT MANUAL

Figure 2-7. Chan A, Typical (Cont)

NOVEMBER 21,1978 Revised March 2,1989

Pan 2 Section I1 Loading Information

fiy r e 2-7.Chan A. Typical (Cont) NOVEMBER 21,1978 Revised March 2,1989

Pan 2 Section 11 Loading information

Aflm~?m7J& FLIGHT MANUAL

Figure 2-7. Chan A. Typical (Cont)

NOVEMBER 2 1.1978 Revised March 2.1989

FUGHT MANUAL

Figure 2-7. C h n A. Typical (Cont)

NOVEMBER 21,1978 Revised March 2,1989

Pan 2 Section I1 Loading information

Pan 2 Section 11 Loading Information

~FUGHT ~ ~MANUAT ~Bs~wIY//E.

Figure 2-7.C han A. Tupi~xl(Con1

NOVEMBER 21,1978 Revised March 2,1989

FLlGHf MANUAL

figure 2-7.Chan A. Typical (Cont 1

NOVEMBER 21,1978 Revised March 2,1989

Pan 2 Section 11 Loading Information

Part 2 Section I1 Loading Information

N G W MANUAL

-

NOVEMBER 21,1978 Revised March 2.1989

fLfGHT MANUAL

AIRCRAFT ACTUAL WEfGHT AND HORQONTAL BALANCE

FOR 5-76 MODEL HEUCOPTER

brpl No.

Reg. No.

i

I

I

I

I

I

LEFI M A I N

1

RIGKT M A I N

f

NOSE

1

I

TOTAL WEIGHT

I

1

I

I

w~ w~ Wm

I

1 I

MAIN ROTOR

. I

I

a

C

2 0

i

K

= 9T

N o a

-

.

~

+ JACK FORWARD POINT

Y = l W

CENTROID

___I

I AIRCRAm HORIZONTAL C.G.

Figure 2-8.

I

w

Chart B (Sheet 1 of 4)

JAC% A n POINT +

Afm?sm/I& R I M MANUAL

PartzSeCtionII Loading &formation

AIRCRAm ACTUAL WEIGHT AND HORIZONTAL BALANCE FOR 5-76 MODEL'HELICOPTER

BY

s8ut-d-

Date

I L

S m b l No.

Reg. No. REAOION WHEEL OR JACK POINT

I

I

RIGHT MNN

!

I

NOSE

1

1

I

I

1

LEFT MAIN

--

-

t

TOTAL WElGUT

I--

Y~

I

%

I

w~

--

-

W

-----+

VERTICAL THRU MAIN ROTOR CENTROID

=

I

1

m

+ FORWARD JACK POINT

JACK POINT +

AIRCRAFT H O R U O N T U C.G.

WEIGHING ON J A C S

W x W N 1

WEIGHING ON WHEELS

S T+L

- -m

= C.G.

+

-

= CG.

CORRECTED WEIGHT AND HORIZONTAL BALANCE

I

ITEMS ADDED O SUBTRACTED

mTu BIILAWCE

1

WEIGH7 (m)

I

WRUONTAL D l S l (ii)CG. TO FWD DATUM


EYPW GROSS

~co-=t=e)

(larscl~~~ia.-r= i

Figuce 2-8.

n

rn

.

F

w

e

t

Chart B (Sheet 2 of 4)

NOVEMBER 21.

1978

Revised JUNE 6, 1980

~

~

~

I

F U M MANUAL

TOTAL

CbartB (Sheet3cb4)

F-2-8.

NO-

21, 1978 Revised JUNE 6, 1980

mwlS ADDED

ITEMS SUBntACTa,

-

TQTAL

I

I

I

I

F i g m ? 2 8 , CbartB(Sheet4af4) NOVEMBER 21, 1978 Revised JUNE 6, 1980

EMPTY WElOHT and BALANCE RECORD

CORIIIIUOUS HIOOIIV W CHIHO(( IH WRUCtURl OR IQUlPMlNI # l l l t I l N O NllOHI

YODIL

#)I111 HO

CHART E!klkNC#

C ¶I1111

CABIN WEIGHT (POUNDS)

C1 ARM = l35.5

Mwmr/loa

IOMEM/lOO

110

l2Q

l23

130 140 150 160 170 180 190 200 210

133 144 154 164 174 185 195 205 W 226

70 80 90

loo

t20 23u

.

246

250

256

260

267

290

3m 310

320 330 3a 350 360 370 380 390

BAGGAGV CARGO

c2

'6

ARM = 1665 MOMENT/ loo

ARM = 235.0 LIOMN/lOO

236

240

270 280

ENGERS

B1 ARM = 1025

51 62 72 82 92 103 113

50 60

'

PILOT AND CO-PILOT/ ENGER

277 287 a 7 3u8 318 328

338 349

400 410 420 430 440 450

F i i 2-10.

Chart E

-

Cockpit and Cabin CompartmeDt Weight and MomentTable@ketIof2)

Part2SectionZ Loading Information

N M MANUAL

PLOT AND

ENGERS

FOUNDS)

c1

c3

ARM = l35.5 MOMENT/l00

ARM = 1 9 7 5 MOMENT/100

623

909 928

637 650 664 678 691 705 718

732 745 759

m

786 799

8U 827 840 854 867 881 894 908 92l

NOTE: L CAWON M U r r

'6 ARM = 235.0 MWEN-r/100

948

968 988 1007 lo27 1047 1067 1086 1106 ll26 1146 1165 1185

a205

= l244

l2a

l2w

l.304 l323 1343

BE TAKEN TO BE SURE THAT ENGER AND/OR

6166A6VCARGO

COMPARTMENT LOADING DOES NOT CAUSE AIRCRAFT MAXIMUMGROSSWEl6HT AND/OR CSi. Ulnm TO BE EXCEEDED. 2 BA66A6VUR60 CENTROD MUST BE LOCATED AT COMPARTMENT CENTROID. 3. MAXIMUM AUOWA8l.E ROOR LOADING FOR BAGGAGE COMPARTMENT IS75 FOUNDS PER SQUARE FOOT FOR WEIGHTS UP TO 600 FOUND YAXlMUYt COMPARTMENT

Figure 2-10. 2-32

Chart E

-

Codpit and cabin Compartment WeMoment Tahle (Sheet 2 of 2) NO21, 1978 aevised JANUARY 9, 1979

a@

- --

N M MANUAL

r

FUEL s Y m M

WEIGHT-LB

NOTES: 1.

2.

MOMENT/lOO

TOTAL

-

1

2 TANKS

'

TOTAL WEIGHT-LB

I

MOMENT/loo

I

THE SINGLE ASTERISK INDICATESTHE APPROXIMATE WEIGHT AND MOMENT1100 FOR F U U FUEL TANKS USING JPP FUEL AT 6 5 POUNDS PER GAUON UNDER nANDARD CONDITIONS (60-F).

(')

THE DOUBLE &K 1NDlCAE.STHE APPROXIMATE WElGW AND MOMENT/lOO FOR F U U FUEL TANKS USING JP-1NELAT6.7SPOUNDSPERGALlON UNDERSTANDARD CONDITIONS (Men. (")

THE TRIPLE ASTERISK INDICATES THE APPROXIMATE WEIGHT AND MOMENT/lOO FOR FULL FUEL TANKS USING JP-5 NUAT 6.8 POUNDS PER GAUON UNDER STANDARD CONDmOW (WF).

3. fa")

4. THE TOTAL USA8t.E CAPAUTY OF 28l2 U.S 6BWOWS (140.6 GALLON5 PER TANK) BASED ON ACTUAL TEST RESULTS.

5. THE USABLE FUR TABLE (6RAVfTY NEUNG) AND EQUAL WEl6im OF FUEL I N EACH TANK

IS

THE FUEL ARM TABLE ARE FOR

6. THE STANDARD 576 HELICOPTER IS EQUIPPEDWITH A CAPACITOR TYPE FUEL QUANTITY

GAGE WHICH IS CAUBRnfED I N POUNDS. AlRCRACTNUVARlE3 I N WEIGHT PER GAUON DEPENDENT UPON T HESPECfnCGRAVTTrAND TEMPERATUREOFTnEF U E L 1HEREFORE. THE NOTATION DOES NOT APPEAR ON THE INSTRUMENT DIAL THE PILOT SHOULD ANTICIPATE v w n o r t s IN THE IWSTRUMENT READINGS WHEN TANKS ARE FUU. 7. THE FUEL ARM T A B E IS INCLUDED FOR B N C E .

FUEL ARM t A 6 l E

1/4 F U U i f 2 FULL 314 FULL F U U (JP-4) FULL (JP-1) FULL (JP-5)

Figure 2-11. Cf;art E

- Usable Fuel Weight and Moment Tahle ( G m m Fueling)

I

ENGlNE OlL TABLE

-

ENGME OIL SYSTEM 2 TANKS TOTAL CAPACITY = 254 US. GALLONS (2 TANKS) ARM = 231.0 MIL-L-7808 OIL MIL-L-23699 OIL GALLONSr

-

WEIGHT POUNDS

25.

,

WEIGHT POUNDS

18 36 45

8.4 16.8 21.0

7.7 15.4 l 9 3

1 2

mo OIL WLER~ANKS

-

MOMENT/lOO

MOMENT/100

19 39 49

CONTAIN ~nGALLONS EACH FOR A TOTAL CAPAC~TYOF 2 sGALLONS.

WINDSHIELD WASHER FLUtD TABLES

I1

CAP-

us. QUARTS

= 5 2 US. QUARTS ARM = 59.8

-- -

Iw ~ ~ ~ ~ r - m r m W ~ I ~ / i o o

FOR THE ALTERNATE nUED TANK LOCATION (7620001034)

I

Figare 2-12. Chart E

- EngineMoment O i l and Wrndshield Washer Flaid Weight and Tables

NOVEMBER 21,

1978

Revised 3atp 26, 1982

IWERNAl. CARGO TABLE CABIN (POUNDS) ARM = 134.25 MOMENl/lOO

ARM = 16825 MOMENT/lOO

13 27

134 201 269 336

17 34 50 67 84 101 118 135 l51 168 252 337 421

403

50s

470 537 604 671

589 673

40 54

67 81 94

107

l21

.

m 841

MAXIMUM CAPAClTY OF COMPARTMENT C4

2 CARGO CENTROID MUST BE LOCATED AT COMPARMEM CENTROID. 3. MAXIMUM ALLOWABLE FLOOR LOADlNG FOR CAR60 IS 75 POUNDS PER SQUARE MQT. 4. SEE COCKPIT AND =N WMPARTMENT TABLE FOR 6A-E COMPARlMENT (C6) WEIGHT AND MOMENT/l00 VALUES AND LOADING REQUlRE?dMS. S=(m

NOVEMBER 21, 1918

EXTERNAL CARGO TABLE WEIGHT (POUNDS)

ARM = 1815

WElGHI

mums)

I

W I M U M m N A L CARGO WEIGHT

NOTE: 1-THS TABLE S APPllCABLE ONLY WHEN THE

LOAD SYSTEM OPTION fS

IWSTAUED.

2 CAUTION MUSTBE TAKENTO BESURETHAT--0.AIRCRAFT MAXlMUhl6R03S WEIGHT AND/OR C6. UUKS TO BE n

DOESNOTCAUSE

mm.

5

I

Figpre 2-14.

chart E

- Exkmal Cargo Weight and Moment Table

NOVEl5ER 21, 1978 Revised APRlL 12, 1979

5

0

Part 2 Section II Loading Information

FUGHT MANUAL

SUDlUG DOOR TABLE

UTUllY INTERIOR:

58 58

FORWARD (CLOSED) AFT (OPEN)

99

lt2

VIP INTERlOR FORWARD (UOSED) AFT (OPEN)

NOTE: 1. THU TABLE U APPLICABLE ONLY WHEN THE SLIDING DOOR OPTlON IS INSTAUEO

s

F v e 2-14A. Sliding Door Weight and Moment Table

NOVEMBER 21, 1978 Revised JUNE 6, 1980

m (Cg)

I

S-70 SAMPLE LOADINQ CONDlTlONS

ER :R MISSION ITEM

WEIOHT (POUNDS)

C A R 0 0 MISSION

MOMENT/lOO (IN-POUNDS)

MOMENT/lOO IIN.POUNb9)

EMPTY WEIOHT PILOT COPILOT ENGINE OIL WASHER f LUlD 1 7 0 18,) fORWARO ENOERS (3 A f t ENOERS (2 @ 170%.) BAOOAOE FUEL UJABLE FLOOR CAROO, FORWARD f LOOR CAROO, A f T

5625 170 150 19

OROSS WEIOHT rnn~.orr; EXTERNAL LOAD

8075

.

EXTERNAL LOAD MISSION WEIOHT (POUNDS)

11

1000 300 BOO

OR099 WEIQMT HOVER-START' MINUS WASHER FLUID MINUS FUEL.USA0LE PLUS RESERVE/LANOINO FUEL MINUS EUTERNAL LOAD OROSS WEIOHT LANDINO'

NOTE: I. SAMPLE LOADINO CONDITIONS DEtERMlNL ACTUAL LOADINO TOR EACII FLIOHT U S I N 0 LATEST WElOHT AND MOMENT/100 DATA. 2. FROM C.O. CNART. FIOURE 2.1. THE AIRCRAFT WEIOHT AND MOMENT/lOOIS WIT11115 LIMITS BALANCE. AND THERE~OREIN

S 53559 ( R I )

LOADING INX)RMATION

Each copy of this supplement which is assigned to a specific aircraft (i.e. aircraft c m ) mast contain all data referred to herein and must be m a i m e d to the current aircraft configuration. Copies issaed on general distribution only need not contain such daQ

NOTE Use this section for helicopters carrying a load on the utility hoist. Use Section 11 for helicopters without a u m t y hoist or for helicopters with a utility hoist installed but no hoist load

The current empty weighf horizontal rno&/100, lateal moment/100, a m center of location horizantal center of gravity location and t are shown in the E m m Weight and Balance Record, Chart C herein. 1.

The cruzent empty weight inelades all items marked (4 in the Empty Weight Check List, Cbart A, herein. Items not included are marked (0)

2.

-

Clmrt C (Figure 3-11) must be kept up to date. If anp items of equip ment are added to or removed from the aircraft, Charts A and C must be updated as follows:

a When W are added, make the entry in the EmpQ Weight and Bahace Record, Cfrazt C. On ckart A mark (4,in both the *TnAiztraft.' and "Chart C Entry"colunms opposite the items added, dalhg the cohnnns where m t e d . Add the new ~orentriesonChartCto~pre~~talsofemptpw horizmtal moment/lOO, and latealmoment/lOO and compute the new emptp weight, horiPrntal moment/100, lateal moment/100, brizontal center of gravity location aad laterdl center of gravity location.

b.

Whenitemsareremovedframthe~foltawasimitarprocedure to (a)above except mark (0) in .the "In Aircraft" colmrm andmask (0) inthewChartC Ensg"colmnnof C k u t A , andsubtract the new entry or eniries from the previous totals of empty weight, brizontal moment/lOO and lateral moment/lOO, sfrown on Chart C (Figure 3- 11).

Part 2 Section IE Hoist Loading Wormation

NCHT MANUAL

New items to be added to the aircraft and not previous1y listed, may be "tmitten in" on the next available line m the appropriate aircraft compartment. Do not erase o r obliterate aay Chart A entry when items a r e removed from the aircraft.

The a i r c a f t empty weight center of gravity location both horizontal and latew is initially obtained by actrral weighing (Figure 3-10). The updated horizontal center of gavity location is determined by dividing the updated horizontal moment/100 by the updated empty weight and multiplying the result by 100. The updated laterdl center of gravity is obtained by dividing the updated l a t e a l moment/ 100 by the mrilafed empty weight and multiplying the result by 100. When determining the updated lateral moment and m t e d lateral center of gravity the correct sign convention must be used. For

items located on the port side of the helicopter longifadinat centerline (left-kmd side facing forrpanl) the late- mornent/100 and corresponding lateal center of gavity a r e positive (+). For items located on the skboard side of the helicopter longitndinal centerline (right-hand side facing forward) tbe lateral moment/100 and corresponding lateral center of gravity are newtive (-). The "In Aircraft" colnmn in Chart A may be used for periodic checking of the *raft inventory. It is recommended that the first check be made prior to the first flight by a new ope-.

-

' I and lateral center o f gravity lodions The original e w weight, h were determined by SikAirtraft, as shown in the Aircraft Actnal Weight and Balance Form Cbart B (Figure 3-10) which is inchded in the airtraft copy of this manual for reference.

Any

changes to charts A, B, or C mast be made in both Sections 11 and m n if-

and the ntillty hoist loading supple-

empty weight, aircraft empty weight Obtain tbe rrent horizontal moment/100 and aircraft emptp weight lateral monsent/lOo from Chart C (Mgrrre 3-11) and enter each in the -te c o w of the manifest.

Obtain the aircraft t a k e 4 gross weight, horizontal and lateral centers of gravity by adding to the airua£tempty weight, horizontal moment/100 and lateral mornent/100 f o r the weight, horizontal moment/100 and lateral moment/lOO of eaEh useful load item to be tarried. ' h e weight, horkmtal mo-t/100 and lateral =menth00 of each usefnl load item are pzovided in Chart E Weight and Moment C-. 'Usefnl load items indude fix& oil, crew, engers, uUity hoist load, baggage and/or cargo. Do not exceed the specified compWment at any time. '

Determine by reference to the -Horizontal Center d GravitJI W (Figure 3-1) and the Lateral Center of GraviQ Chart NOVEMBER 21, 1978 Revised JUNE 6, 1980

(Figare 3-31, m t the take-off gross weight, horizantal moment/ 100 and lateral moment/lW are within the recommended weight and horizontal center of gravity and lateral center of gravity Limits. The aft horizantd center of gram limit has been adjusted to a c c m t for the landing gear horizan&l retraction moment/100 of plus (+) 33. Therefore, loadings based on wheels d m conditions which fdl within the limiting horizontal moments/lM) of Figme 3-1 wiU be within horizontal center of gravie limits for flight with the fanding gear r e m t e d . Landing gear r e m t i a n has no effect on lateral center of graoiq. 4.

FOR AlRCRAFI' WITH THE UTILITY BOIST OPTIOK mSTALLED AND OPERATED DUEUNG FLIGHT, CHECK THAT TEE AZRCRAm WIU REMAZN WITBIN WEIGHT &D CG W S THROUGBOUT TEE FUGHT: (a) Compute the infiight and &ding gross weight horizontal moment/l00 and lateral moment/100 for landing, and each h£li&t use of the utilie hoist considering foe1 bum-off, sliding door operatian, hoist operator mwemcrat, and any useful load added or removed during the flight.

-

(b) Using the Allowable Hoist Load (Figore 3-4) determine the &owable hoist h d for each use of the

umity hoiin

(c) Determfne by reference to the IIoxhmW Center of Gravity Chart (Figare 3-1) and Lakal Center of Glachart (Figure 3-3) that the airrraft w i l l temain within the recommended weight, horfio4tal center of g r a mlimfts and lateral center of gravttp limits Wfth and wlfhout a hoist load Bring tbe aircraft within weight, hoand lateral k b a c e if either is outside the recommended 'iimitfi by re&c&g allowablehoist Load and/or redistribrrtion of asload '-

NOVEMBER 21, 1978 Revised OCTOBEZt 14, 1980

Horizontal Center of Gavity Chart

NOVEMBER 21, 1978 Revised MAY 14, 1985

~ E a,1978R Revised MAY 14, 1985

LATERAL CEMTER OF QRAVITY CHART

(3oosrtrvri m o m t ~ t / l o o( r ~ . o o u n o s )

41. NEOATIVE

MOMENT/IOO (IN.POUNDS)

I 3-8

Figure 3-4. Permissible Hoist &ad for Combinations of Gross Weight and fltP_ral Center of Gravity

NOVEMBER 21, 1978 &vised MAY 14, 1985

REF UNE MAIN ROTOR

LEVEL AIRCRAn AT STA 176

CENTROID MAIN ROTOR

Q MAIN ROTOR S

-

200 7

H m BAGGAGE

I

I 1628'

500

550

DIMENSIONS FOR WElGUtNG AIRCRAFF ON WHEEL5

DIMENSONS FOR WEIGHING ON JAW5

AIR-

NOTE

COMPIRTMerrS

A ElEGfRON1CS COMAPRTMEKT

AU DtMENSlOUS ARE APPROXIMATE

B wCI(Prr

DIMENSIONS FOR WEIGHING ON WUEElS (J-LRT) MUST BE DmRIdlNEO BY ACTUAL MEASUREMrn-

C CABIN

D MWNE AND TRLLNSMlSSlON COMPARTMENT E TAIL f -EQUIPMENT

YUN LG WHEEL Q AT B L 49.4

W N LG W H E n

ON -6E

DIMEN9W S FOR WElWlN6 AmcRFrONJlSQCS

DIMENSONS FOR

WEIGHJNG H E L I W m ONwHEElS

BL ( ) ,

L.4 I

I

3Lb.4 (+I

Figure 3-5. Aircraft Diagram NOVEMBER 21, 1978 Revised JUNE 6, 1980

WElGHlWG PROCEDURE 1. THE AIRCRAFT MUST BE LEVU HORQONTAUY AND U l E R A L l Y WHEN WEiGHED. !T IS RECOMMENDED THAT THE AIRCRAFT BE WEIGHW ON JACK POINTS TO PROVIDE EASE I N W N G . 2 SUSPEND A PLUMB B b B ABOVE THE LML PLATE (LOCATED ON THE HAND ENGER DURING WEIGHING DOOR FRAlltE AT STATION 176) FOR LEVEUNG THE AIR= 3. WHEN WEIGHING ON JACK POINTS. ACTUAL MEASUREMENTS DURING WEIGHING ARE NOT NECESSARY. *mR W N G THE AIRCRAFT THE FOUOWNG DIMENSIONS U Y BE ADDED DIRECTLY TO THE WEIGHING FORM (CHART B): K (HORIZONTAL DATUM T0

%

OF FORWARD JACK POINTS)

N (HORIZONTAL DATUM TO Ci OF AFT JACK P O I m

=

9 2 0 INCHES

= 252OINCHEs

4. WHEN WElGHlNG ON WHEELS. MEASURE DIMENSIONS J.LRT. AFER L M U N G AND DURING WEIGHING. USING THESE ACTUAL DIMENSIONS AND THE FORWARD JACK POINT (AIRCRAFf S7ATtON 9 D 2B .0).l M T U t Y t N E THE DISTANCE FROM THE HORaOMAL DATUM TO THE 5 OF THE LATERAL DATUM TO THE G-OFTHEMAIN WHEELSUSING BOTH FORWARDAND m JAC)(POIM(SEELATERAl C G MEASUREMENT PROCEDURE).

FOR U3ECKWG PURPOSES THESE DIMENSTONS ARE APPROXIMANY AS FOLLOW%

IWRQONTAL

DATUM TO THE

-ALDATUMTOTHE LATtRALDAfUYl707HE

q

LATERAL DATUM TOT HE^

$ OF THE NOSE WEEL

=

q OFTHEYAtNrnmLS

= 2W.O

OFTHENOSEWHEU

=

OlNCHES

OFTHELEFT W N W E E L

=

49.4 INCHES

=

49.4 INCHES

L A T E R A L D A W T O T H E q WllfEM6HT W N W H E E L

60.OIUCHES INCHES

N U E WHEN WBGHINC ON WHDLS.THE W N ANDNOSEOLEOS SHOULDBE ~ ~ l l ff a f PROPER ~ ~ ~ FLUID m LEYELS m~ ~ R I U L Inuto C r~ THE

Figure 3-6. Weighing Procedure (Sheet 1 of 3) NOVEMBER Zl, 1978 -Vised JUNE 6,1980

Part 2 Section m

SA 4047-76-1

FLIGHT MANUAL LATERAL CENTER OF GRAWrY MEASUREMENT

PROCEDURE FOR WEIGHING ON UlEEU TO CALCULATE THE AIRCRAFl LATERAL CENTER OF GRAVITY THE DISTANCE FROM THE NOSE AND MAIN WHEELS FROM LATERAL DATUM (BL 0) MUST B E MEASURED. BL 0 OF THE AIRCRAFT CAN BE DEERMINED FROM MEASUREMENT BY SUSPENDING A PLUMB BOB FROM THE THREE JACKING POINTS AT STA 92. BLO. STA 252 BL 25 (LHS). BL 25 @+IS). MEASURE CENTER O F THE A m JACKING POINTS AT STA 2S2 AND THIS POINT LONGlfLIDlNAUY WITH THE PMNT AT STA 9 2 T H I S WILL SET U P A LONGITUDINAL U N E AT BL 0. THE DISTANCE FROMTHE 5 OF THE LEFT AND R l G H l MAIN WHEELS CAN THEN BE MEASURED FROM THIS BL 0 UNE.

LATERAL DATUM BL 0

BL -25RHS (STBD)

RH MAIN G l B D )

WHEEL

LH

WHEEL

PORT)

Figure 3-6. We-

Procedure (Sheet 3 of 3)

NOVEMBER 21, 1978 Rtvised JUNE 6, 1980

CREW.

ENGER,AND BAGGAGE COMPARTMENT CENTROIDS

AVIONICS CDMPARTMEHT

C

CREWENGER CErnID

-GAGE

NET)

~ M P A R T -WEIGHT HORQONTAL LATERAL MENT OUNDS) CENTROID CENTROID

COCKPIT B1

Bl

1m 170

11-

5 % %

17D YO 17D .17D 170 170

YlEM

+Y,O

%?

-18.0

=2

u5.5 u5.5

+27&

US5

-9.0

c, % c, %

l.355

-274

BASGA6E

1665 1665

+27&

USIN

cl Cl

WMPACCF- WEIGHT HOIUZOHTAL UTERAL

*&

*.0

c6 c6

(POUNDS)

SfROID ITRDID'

l70 170

1665 1665

-2?A

m

ms

+W&

170 170

*QO

M

l975 l975 IS75

-27.0

300 300

2350 &O

+l83 -I83

Figure 3-7. Standard 14 Place -C enger and Baggage Centroids NOVEMBER 21, 1978 Rmiaed JUNE 6, 1980

m,

-9.0

-9.0

s5Tmm

'

1 3-13

CARGO C O M P M E N T CENTROtDS

BASED ON THE MAXIMUM ALLOWABLE FLOOR LOADING OF E POUNDS PER SQUARE FOOT. THE W M U M W A G m OF THE FLOOR BEmEEN FLOOR 'IIEWWN R l W W A D E D AREA) tS 1300 POUNDS.THE CMTROU) OF THE SHADED AREA 1S STATION 15025.

AVIONICS COMPARTMENT

= m C U )

WOPOUND SIRUCTURU

w m o

.puxuIUUALLOIABLfFLOOR UMMWGm~mIs75 KHmwPERsQuAREFoar YUIMUY ULOIVA8LEfLmR LOADlNC FOR Bh6me

COYPARlMPnRbORLS7s ~FERSQUAREFOOTmR

WEl6HTS UP TO THE y.*lyBA6GA6E w M P A R T Y P ( T ~ H T W 6 0 0 P O ~

3-

1 3-14

F i e 3-8. Standard Cargo Ca-n NOVEMBER 21, 1918 Rsvissd J7JN5 6, 1980

- Cargo Centroids

FLIGHT MANUAL

Figure 3-9. ChartA

NO-

21, 1978

Revised JUNE 6, l980

Figure 3-9. Chart A (Contispgd) NO?%= 21, 1978 Revised JUNE 6, 1980

WGHT MANUAL

r i g c u e 3-9. c%artA (Continued) NOVEMBER 21, 1978 Revised JUNE 6, 1980

Figure 3-9. Chart A (Continoed) NOVEM&R 21, 1978 Revised JUNE 6, 1980

R l M MANUAL

Figure 3-9. Chart A (Contimed) NOVEMBER 21, 1978 &Vised JUNE 6, 1980

mGHf MANUAL

Figure 3-9. Chart A (CantirrPed) NOVEMBER 21, 1978 Revised 3une 6, 1980

Figure 3-9.

Chart A ( C a u e d )

NOVEMBER 21, 1978 EkvbedJUNE6, I980

Figure 3-9. Chart A (Continued) NOVEMBER 21, 1978 Revised JUNE 6, 1980

FLIGHT MANUAL

Figure 3-9. Chart A ( C o n m d )

F

i 3-9. C b r t A (Co-d) NOVEMBER 21, 1978 Revised JTME 6, 1980

F U M MANUAL

Figure 3-9. Chart A (Continued) NOVEMBER 21, 1978 &!vised JUNE 6, l980

FUGHT MANUAL

Figure 3-9. Chart A (Continued) NOVEMBER21, 1978 Rb-d JUNE 6, 1980

Chart A (Cantbed) NOVEbdBER 21, 1978 Revtsed JUNE 6, 1980

Figare 3-9.

FUGWf MANUAL

AIRCRAFT ACTUAL WEIGHT AND HORIZONTAL BALANCE FOR 5-76 MODEL HELICOPTER

S H E E T OF-

PREPARED BY DATE REG. NO.

t

t

I

LEFf MAIN

I

RIGHT MAIN NOSE

( :!

TOTAL WEIGHT

i

SERIAL NO.

1

i

1

WR

!

i

WN

I 1

I

!

w

!

!

.

I

i

I + FORWARD JAcKpO'M

I

AlRCRACT HORIZONTAL C G WEIGHING ON MQtS WEIGHING ON WHEELS

Figme 3- 10. Chart B (Sheet 1 of 6) NOVEMBER 2 1, 1978 Revised JUNE 6, 1980

A n JACK POINT

I

I

A

1

+

UTERAL DATUM VERTICAL THRU MAIN

ROTOR CENTROID

RHS (STARBOARD) JACK POINT

+

I

l.Hs 0

JACK POINT .c

RH 'MAIN

W MAIN WHEEL

WHEEL

+

I

AlRCRAn U m L CG. NOTE NEG4lWE CG= W G H l SIDE (STARBOARD)

NLlD JACK

AIRCRAFT ACTUAL WEI6Hf AND HORIZONTAL BALANCE FOR 5.76 MODEL HELICOPTER PREPARED BY DATE

S H E n O F -

S E R W NO

REG. NO

I

LEFT MAIN

.

I

I

I

RIGHT MAIN NOSE TOTAL WEIGHT

1

1 1

.

.

WL

I

1

1

WR

1

1

Figure 3-10. Cbart B (Sheet 3 of 6)

NOVEMBER 25, 1978 Rsvimd JUNE 6, 1980

WN W

I

I

I I

LllTERIL

DATUM

VERTICAL THRU MAIN ROTOR CENTROID

RHS (STARBOARD) JACKPOINT

I

.

9--+25'

0

I

T-

+

1

+

-4

t,

RH MAIN

WHEEL

,

LHS (PORT)

JACKPOINT

R

-----------I

LH MAIN WHEEL

E l

FWD JACK WlNT

AIRCRAFT LATERAl C G NOTE NEGATIVE C 6 = RIGHT SIDE (STARBOARD) WEIGHING ON J A W W

L x 25)

-

(W

x-25) =C G

W

CORRECTED WwIt AND BALANCE ITEMS ADDED AND

su8mA~

AIRCRAFT llS WEIGHED

TOTAL

WEIGHT

,

HORIZOHTU DISTANCE QN)CGfO

(LB)

FWD DATUM

I

ldoMENT
U-DATUH ON) C G TO

-

- -

*won
1

WEIGHT

GROSS

ChartB (Sheet 4 of 6) NOVEB43ER 21, 1378 Revised JUNE 6, 1980

B

Figure 3-10.

3-35

.

Figtme 3-10. Chart B (mt5 of 6)

NOVEMBER 21, JUNE 6,

-vised

3978 1980

Section ID 4 f l ~ ~ ~ ~ ~ dPartg2Information & FUGHT MANUAL

-KEMS ADDED KEM

WEIGHT

ARM

TOTAL

KEMS SUBTRACTED

TOTAL

I Figure 3-10. Chart B (Sheet 6 of 6) NOVEMBER 2 1, 1978 Rsvised JUNE 6, 1980

MOMENT

Part 2 Section EI Hoist Loadin.- Information

&H@%!!KW~& NGHT MANUAL

*

EMPTY WEIGHT AND BALANCE RECORD COWnNWUS H(ST0F.Y OF M A N G E S IN STRIICTSREOf EQUIPMENT AFFECTING WUGWT

DESCRIPTION OF lRTlCLE OR MODIFICATION

NOTE: POSITIVE LATERAL C G 'S ARE LEFT (WRT) SlDE NEGATlVE LATERAL C G 'S ARE FOR RIGHT (STARBOARD) SlDE

Figure 3-11. Chart C (Sheet 1 of 2) NOVEMBER 2 1, 1978 Revised JUNE 6, 1980

LW C E

Part 2 Section I II

&8ff%!g%? Hoist % Loading E Information FUGHT MANUAL

r

EMPTY WEIGHT AND BAANCE RECORD CONTINUOUS HISTORY OF CHANGES IN STRUCTURE OF IlufpMENT AFFECTING WEIGHT (L BAUNCE

OESCRWTION OF ARIICLE OR MODIFICA?lON

NOTE:

PosmvE LATERAL CGSARE LEFT (PORT) SIDE

NEGATIVE LATERAL CG'S ARE FOR RIGHT (STARBOARD) SIDE

Figure 3-11. (Jhart C (Sheet 2 of 2) NOVEMBER 21, 1978 RavLsed JUNE 6, 1980

-

COCKPIT- B.

-

CABIN ENGERS WEIGH1 POUND!

-

PILOT AND COPILOT/ ENGER HORIZONTAL ARM-1025 MOMENT/10(

PILOT LATERAL ARM= -18.0 MOMENTA0

COPILOT/ ENGER LATERAL ARM= +17.0 MOMENT/lOC

HORIZONTAL

10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250. 260 270 280 290 300.

NOTES: 1. CAUTION MUST BE TAKEN TO BE SURE THAT ENGER AND/OR BAGGAGWCARGO COMPARTMENT LOADING DOES NOT CAUSE AIRCRAFr MAXIMUM GROSS WEIGHT AND/OR C G uMrrs TO BF FXCFFDFD - --2 BAGGAGWCARGO CENTROID MUST BE LOCATED AT COMPARTMENT CENTROID. 3. MAXIMUM ALLOWABLE FLOOR LOADING FOR BAGGAGE COMPARTMENT IS 75 WUNDSPERSQUARE FOOT FOR WEIGHTS UP TO 600 POUND (300 POUND/SIDE) MAXIMUM COMPARTMENT WEIGHT. 4. LATERAL ENGER C G DATA IS TO BE CONSIDERED FOR EACH SEAT ROW (C1. Cz C3) 5. LATERAL C.G. DATA L E IS ~m s r n V E (+I RIGHT IS NEGATIVE (-1. *MAXIMUM CAPACITY OF COMPARTMENT C6 (PER SIDE)

--------

Figure 3-12.

-

Chart E Cockpit and Cabin Compartment Weight and Moment Table

NOVEMBER 2 1, 1978 Revised OCTOBER 14, 1980

BAGGAGWCARGO C6

1

LATERAL

LATERAL

ARM = -9.0 MOMENT/lO(

ARM=, +I83 MOMENT/10(

LATERAL ARM= -183 MOMENT/10(

RIGHT MANUAL

-

-

f UEL SYSTEM 2 TANKS CAPACTTY=

-* I

TOTAL 10Gm

M0MENl/100

2 8 UW N S

fANI(S)

TOTAL YPEIGHT LB

-

No1. f) THE SINGLIE ASTERlSK INDICATES THE APPROXIMATE WEIGHT AND YOMENT/lOO FOR f U U FUEL TANKS USING J P 4 FUEL AT 6 5 POUNDS PER GALLON UNDER STANDARD COUDmONS MWn. 2 pg)THE DdUijL€KINDICATES THE APPROXIMATE WEIGHT AND M O Y f M / 1 0 0 FOR F U U NUTANKS USING JP-1 FUEL AT 6.75 POUNDS PER GULON UNDER STANDARD ~ D r n O U l6nm. S 3-F%%E ~ E R E INDlCATES K THE APFROXIYATE WDGHT AND MOMEHf/100 FOR F U U F U R TANKS USING JP-5 FUEL AT 6.8 POUNDS PER GAUDN UNDER STANDARD CoUDrnONs (6ov)4. THE TOTAL USABLE CAPACTTY OF 2BLZ US. GAIlI)NS (140.6 GALLONS PER TANK) IS BASED ON ACTUAL TEST RESULTS. 5. THE USABLE ML TA3I.E (GRAVllY N E U W G ) AND THE NU ARLll TABLE ARE FOR EQUAL WElGHTS OF FUEL I N EACH TANK 6. THE STANDARD 3.76 HELICOPTER S ! EQUIPPED W l l H A UPAQTOR TYPE FUEL Q U m GAGE WHlCH tS CALI%RATED IN PWNDS. AIRfRlVT FUEL VARIES IN WEI6KT PER C*LLDN DEPENDENT UPON THE SPEUFYC GRAVrrY AND EMPERANRE OF THE FUELTHEREFORE THE NOTATION TUU' DOES NOT APPEAR ON THE INSTRUMENT DUL THE PILOT SHOULD ANTIUPATE VARlATIORS W THE INSTRUMENT REAMNGS WHEN TANKS ARE f u u 7. THE FUEL TABLE IS INCUtDm FOR REfEENtL

rrurrukk

mARwIrrstr

Figure 3-13. Chart E

-

Usable Fuel Weight and Moment Tabk (Gravity Fueling)(Sheet I of 3)

NOVEMBER 21, 1918 Wviscd JUNE 6, 1980

-

- -

FUEL SYSTEM 2 TANKS CAPACITY= 140.6 GALLONS PER TANK

II

LEFT TANK

I 3-42

Fig~r3 e-l3. Chart E (Gra*

RIGHT TANK

- Usable Weight Moment Table FPeling) (Sheet 2 of 3) ?bd

NOVEMBER21, 1978 F & v h d JUNE 6, 1980

and

FUGW MANUAL

NOTES: 1. (.) THE SINGLE UfERlSK INDlCATn W E APPROXIWE WE16HT POUNDS AND MOMENT/ 100 FOR FULL FUEL TANKS USlNG JP4N AT PER GALLON UNDER STANDARD CONDmONS (WAND 2 pm) THE DOUBLE ASrauSK INDICATES 'IHE APPROXIMA= MOMEN7/100 FOR FULL FUEL TANKS USING JP-1NU AT 6-75 POUNDS PER GALLON UNDER STANDARD CONDmOHS (600F). 3. POm) ME TRIPLE AmRfSK INOtCAfES W E APPROXIMATE WEIGHT AND MOMENT/ 100 FOR FULL FUEL TANKS USING JP-5 FUEL AT 6.8 POUNDS PER GALLON UNDER STANDARD CONDm0N.S (WF). 4. THE TOT& USllBLE CAPACTPI OF 283.2 U S W O N S 040.6 GALLONS PER TANK) IS BASED ON ACTUAL TEST RESULff. 5. THE STANDARD 3.76 HUlCOPTrR IS EQUiPPED m H A CAPAQTOR-rn FUEL QUANTITY GAGE WHICH IS CALIBRATED IN POUNDS. AlRCRAfl NU VARIES I N WE1GW PER G W O N DEPENDENt UPON THE SP&lFIC GRAVITY AND TEMPERATURE OF THE FUEL THEREFORE. THE NOTAflON WU' DOES NOT APPEAR ON THE INSfRUMENT D l A L THE PILOT SHOULD AKnClPA VARIATIONS tN THE INSTRUMENT READINGS W E N TANG ARE NU 6. THE FUEL ARM TABLE LS INCLUDED FOR REFERENCE. 7. LATERAL C 6 DATA EFf If PMmVE (+),RIGHT IS NEGATIVE (-).

-

FUEL SYSTEM 2 TANKS

I

I

PER TANK QUANmv

)(OWZOMAL

PER TANK WElGKT

ARM

LB

I

MOMENl/lOD

EFf LATUW ARM

I

MWMflOO 1

Lf4FUU 1/2 F U U 314 F U U

Figure 3-13. Chart E

- Usable Rrel WeigirtandMoment Table

(Gravity Fueling) (Sheet 3 of 3)

NOVEMBER 21,

1978

Revised JUNE 6, 1980

I

Part 2 Section III H O L ~~mm ~ xaformatiop

AFlUmM MANUAL ~~mdE

-

ENGINE OIL TABLE -

-

ENGINE O U SYSTEM 2 TANKS

254 US. GAUONS (2 TANKS)

TOTAL CAP-=

1

WL-L-7808 OIL GALLONS

(mu-)

Figure 3-14.

HOWZONTAL

LATERAL

ARM= 2310 MOMPCT/lW

ARM= +10.0 MOMPJT/lOo

Chart E

- Engine Oil Weight and Moment Table

WINDSHIELD WASHER FLUID TABLES

CAPAUIY 5 2 US. QUARYS

LATERAL

I

ARM= +17B YOldEMAOo

t

I

/ Fif3n-e 3-15. 3-44

CAPluxn 4 Us. QUARTS

-

I

Chart E W-eld Washer Fluid Weightand Moment Table NOVEMBER 2 1, 1978 Revised JANUARY 17, 1983

.

WGHT MANUAL

EXTERNAL CARGO TABLE

WEIGHT (POUNDS)

NOVEMBER21, 1978 ReFlised JUNE 6, 1980

Part 2 Section IU S~~RB&& Eioist Loading Ipfosmaticm FLJGHT MANUAl -

UTUfTY

HOIST LOAD TIBLE

*MAXIMUM VllLlTY HOIST LOAO

L THIS TABLE IS APPlKA8l.E ONLY WHEN THE UTILITY H O S l OPTION IS INSTILLED. 2 UCmW MUST BE TAKEN TO BE SURE THAT LmUrY H O W OPERATlON DOES NOT CAUSE A J R C W MAXIMUM 6ROSS WEIGHT AND/OR HORIZONTAL AND LATERAL CENTER OF 6R&VITYUYKS TO BE EXCEEDED.

3. LATERAL C 6 DATA RIGHT IS NEGATIVE 1-1

5-IC9,

-

Figure 3-18- Chart E Utility Eoht Load Weight and Moment Table

NOVEb5ER 21, 1978 Revised JUNE 6, 1980

I

SLIDlFlG DOOR TABLE

I

I

VIP INTERIOR

NOTE: 1 THIS TABLE IS APPLICABLE ONLY WHEN THE SLIDING DOOR OPTION(M0DIFtED BY P/N 760m200l5411)t~INSTALLED

Figure 3-19. Chart E

s-m)

- Sliding Door Weight and Moment Table

HOIST OPERATOR WElGHT AND MOMENT TABLE

HORZONTAL

ARM= 166.0

MOYIENTf loo

.-

+17 +33 4 0 66

50

+83

60 70 80

llo

+lo0 +I16 +l33 +I49 +I66 +l83

I20

+lS9

90

100 230

+216 +232

140

Eo

+249 +266

160 170 180

+a2

rn

200 a0 240

-8

-11

-E el9 -23 -27 -30 39 -38

41 46 49 43 -57 -61 -65

43

+3l5 +332

-72 -76

+365

23Q

4

+299

+a

Po

UTERAL

ARM= 3g0 MOMENTflOO

80

+382

-84 47

+398

-91

4s

-95

1 TMS TABLE

7HE H O m OPERATOR AT THE DOOR STAflOKFOR SEATED VALUE SEE RGURE 8l2

2 LATERAL C 6 DATA WGHT IS NE6AlIVE (-1

Figure 3-20. Chart E

s 9810 cs)

- Hoist Operator Weight and Moment Table

NOVEMBER 21, 1978 Rsvissd JUNE 6, 1980

FLIGHT MANUAL

SAMPLE UTILITY HOIST BAIANCE CALCULATION

ITEM

EMPTY WEIGHT

WEIGHT POUNDS

UITEPa

ARM-

6370 170 170 20 11 60 1828 170

210.4 1025 1025 231.0 528 235.0 235 216.7 197s

13402 174 174 46 6 118 141 3962 336

-05 -18.0 17.0 0 17.0 18.3 -18.3 0 +9a

8849

2075

18359

-02

50

GROSS WEIGHT - TAKEOFF

-

HORIZONTAL

-32

I

COMMENTS

-11 0 +15

FROM CHART C FROM FIGURE 3-12 FROM FIGURE 3-12 LH AND RH VALUES COMBINED FROM flGURE 2A-14 FROM FIGURE 3-15 + FROM FIGURE 3-I2 FROM FlGURE 3-12 LH AND RH VALUES COMBINED FROM flGURE 3 A - U CABIN SEAT LOCATION - FROM flGURE 3A-32

-19

CHECK WEIGHT AND C.G.S '

-31

+29

0 +2 +9

-

ARE WITHIN LIMITS

FROM ABOVE FROM FIGURE 3-20 FROM FIGURE 3-19 FROM FIGURE 3-19 DIFFERENCES BETWEEN TAKEOFF AND MISSION POINT FUELS - FROM FIGURE 3A-13 FROM ABOVE CHECK WEIGHT AND C G 'S ARE WITHIN UMITS FROM FIGURE 3-18 CHECK WEIGHT AND C 6 'S ARE WITHIN LIMITS FROM ABOVE LOCATE LOAD I N CABIN FROM FIGURE 3-16 CHECK WEIGHT AND C 6 '5 ARE WITHIN UMITS FROM FIGURE 3-18 CHUX WUGN AND c G 'S ARE WITHIN U M ~ FROM ABOVE LOCATE LOAD IN CI\BIN FROM FIGURE 816 CHECK WEIGHT AND C G T AR€ WITHIN LIMITS . FROM FIGURE - - - -3-18 --

I

I I

-

I

-

CHECK WEIGHT AND C G 'S

FROM ABOYE

I

ARE W r r n l N LIMITS

-

I

L O C A T I L O ~ N GCABIN FROM flGURE 3-16

c m a WEIGHTANDCG~ w m i

t U ~ M ~ DIFFERENCE BEWEEN MlSSlON POINT FUELS FIGURE 3-13 CHECK WEIGHT AND C G 'S ARE WlTHIN LIMITS FROM ABOVE FROM ABOVE CHECK WEI6HT AND C G 5 ARE W l l H I N UMITS FROM ABOVE CHECK WEIGHT AND C G 5 ARE llVlTHlN UMrrS FROM ABOVE FROM ABOVE

CHECK wii~WrAN~CsAREWITHIN U FROM AR -W-F- .. .

M

I

~

--

%ECK WEIGHT AND C G 'S ARE WITHIN LIMITS

-

GROSS WEIGHT LANDING

FROM ABOVE FROM ABOVE %OM ABOVE X O M ABOVE ?EMOW ALL REMAINING FUEL FROM flGURE 3-13 W D RESERVVLANDING FUEL FROM FIGURE 3-13

6

AL

--

7660

203.5

Figure 3 -2 1. Sample Zloading

NOVEMBER 2 1, 1918 Revised OCTOBER 14, 1980

15585

+O-3

+i !O

YECK WEIGHT AND CGT ARE H I T H I N UMlTS

1

NOTES 1. SAMPLE LOADING CONDITION.DETERMINE ACTUAL LOADING FOR EACH FLIGHT USING UIrrST WEIGHT AND MOMENT/100 DATA. 2 FROM C.G. CHARTS FIGURE 3-1AND 3-37HEAIRCRAFT WEIGHT HORLZONTAL MOMENT/lM) AND LATERAL MOMENT/lW) ARE WITHIN UMlTS AND THEREFORE I N BALANCE 3 EFFECT ON GROSS WElGM OF FUEL BURN O f f MUST BE CONSIDERED WHEN DETERMINING THE MAXIMUM HOIST LOAD M A T CAN BE APPLIED WHEN LOADING OR UNLOADING USING THE LmLrrY HOIST. 4. CARE SHOULD BE TAKEN TO ENSURE THAT THE CORRECT SIGN CONVENTION IS USED. SUBTRACTlON OF A MINUS MWEHT WILL CHANGE THE MOMEHT SIGN f~ msmE MAXIMUM ALLOWABLE HOIST LOAD FOR THIS GROSS W U G M CONDITION IS 490 POUNDS. SEE EXAMPLE ON FlGURE 34.

A FLIGHT H ~MANUAL BEM~~E

Part 2, Section IV Noise Abatement

GENERAL PROCEDURES RECOMMENDED TO MINIMIZE ACOUSTICAL IMPACT 1. In general, maximum distancelaltitude separation from noise sensitive areas is the most effective noise abatement procedure. 2.

Control inputs should be made in as gradual a manner as possible.

3.

Noise exposure is lower upwind from the helicopter than downwind.

4.

Noise exposure is lower on the inside of a turn than on the outside.

5. Noise exposure is lower during right turns than left turns.

Approach Procedure 1. Plan routes and approach paths to keep noise sensitive areas to the left side of the helicopter. 2. Maintain airspeed at 80 KlAS or greater as long as possible. Avoid low speed low rate-of-descent (creeping) approaches.

3. Descend at 85 KlAS and 1200 to 1400 fpm rate of descent (ROD) to minimize noise and maximize separation from noise-sensitive areas.

4.

If more moderate approach angles are required, descend at 95 KIAS and 700 to 900 fpm ROD or 105 KlAS at 500 to 700 fpm ROD.

5.

Delay deceleration to touchdown speed or hover as long as practical.

6. When clear of the noise sensitive area, return to Category A or Category B flight procedures as applicable.

Take-Off Procedure 1. If possible, plan takeoff path away from noise sensitive areas. 2.

Keep noise sensitive areas to the right side of the helicopter.

3. If possible, takeoff into the wind to maximize climb angle.

NOVEMBER 21,1978 Revised November 4,2002

Part 2, Section IV Noise Abatement

I

4.

I

hiRflE~~mflA FLIGHT MANUAL

Avoid right hand crosswinds as additional tail rotor torque can significantly increase climb noise levels. In general, avoiding right hand crosswinds takes precedence over keeping noise sensitive areas to the right side of the helicopter.

5. Climb to cruise altitude at takeoff power and at an airspeed lower than best rate of climb (approximately 65 to 70 KIAS) to achieve a higher climb angle.

Enroute Procedure Maintain a minimum altitude of 2000 feet when transiting noise sensitive areas. When 2000 feet cannot be maintained, reduce airspeed to not exceed 120 KIAS. Plan routes to keep noise sensitive areas to the right side of the helicopter. If noise complaints arise, altitude should be increased.

The procedures described in this section are to be considered as guidelines and should not take precedence over the requirements of normal operating procedures, Part 1 Section II, of the flight manual.

NOVEMBER 21,1978 Revised November 4,2002

FUGtfT MANUAL

~erf&e

Data

SUPPLEMENTAL PERFORMANCE DATA

This section provides p e r f o m e data necessary to compote fuel requiremeats for IT% and VFR fligM in compliaace with FAR Part 91 mles. n e s e data are m e d on daal engine apeMan with Detroit Diesel Allison 250430 and 250430s gas W i n e engins

NOVEBdBER 21, 1978

Revised MARCH 30, 1983

Part 2 Section V Srmb1eme --=- --

FLIGHT MANUAL

Performance Data

LIST OF PERFORMANCE CHARTS

TITLE

PAGE

-

Ihel Flow Hover in Ground Effect at 5-foot Wheel Height at 100%N r ...... Ihel Flow - Hover in Grormd Effect at 5-foot Wheel Height at 107% N r ...... F'uel Flow - Hover out of Gramrd Effect at 100-foot Wheel Iieight at 100% Nr Feel Flow - Hover out of Ground Effect at 100-foot Wheel Height at 10% Nr .... Climb - Dual Engine Normal Cruise Power at 100% N r - Time to Climb and Fuel Co~lsrrmed .................. Fuel C o s m q t i m per Nautical Mile at 100% Nr for Sea Level Aikitpde ........ F'uel Consrrmptim per Nautical Mile at 100% Xr for 3,000-foot Pressure Altitude ............................. Fhel Co~spmptionper Naatical Mile at 1& Nr f o r 6,000-foot Pressure Altitpde ............................. Fnel Co-n per N a a t i d Mile at 100% Nr for 9,000-fdot Pressure AltiWe ............................. F'uel Consumption per Nautical Mile at 100% N r for 12,000-foot Pressrae

....

&We...-...........-....-.....-..

l Consrrmption per Naatical Mile at 100%Nr for 15,000-foot Pressure Altitrrde...,.....-.................-Ehel C0-~1 per Nautical Mile at 10% ~r for sea ~ e v eAltmxie l Frrel Consrrmption per Naotical Mile at 107% Nr for 3,000-foot Pressure m-e. Fuel Cm -t per Naotical Mile at 10% N r for 6,000-foot A.essrae Llltitrade. Frrel Consrrmptian per Nautical Mile at 107%N r for 9,000-foot Presure -We. -1 Consrrmptian per Natrtical Mile at 107% N r for l2,OOO-foot Pressure Altitclde. Fuel Co-n per Nantical Mile at 107% Nr for 15, OMCfoot Pressure Lllti-e..... Beadwind and Taitwind m a c e on Fuel and Crzlising Time Requirements Optimmn Cruise Performance Esample Problem .....................

5-3 5-4

5-5

5- 6 5- 7 5-8 5-9 5- 1 0

5-11 5- 12

b

........

5- 13

5-14

............................

5- 15

............................

5-16

............................

5- 17

............................

5- 18

........................ ...... ..........

NOVEMBER 2 1 1978

Revised MARCH 30, 1983

5- 19 5-20 5-21 5-22

Part 2 Section V Supplemental

FLIGHT MANUAL

Performance Data

-

z

-...-. U

"

=P= 5-1 NOVEMBER 21, 1978 Revised MARCH 30, 1983

NOVEMBER 21, 1978 Revised MARCB 30, 1983

Part 2 Section V Supplemental Performance Data

Figrae 5-3 NO21, 1978 Revised MARCH 30, 1983

Figare 5-4 NOVEMBER 21, 1978 Revised MARCH 30, 1983

FUGHT MANUAL

-

576

C U M B 2 ENGINE NORMAL CRUISE POWER lOO%NR BEST RATE OF C U M B S P W LANDING GEAR UP VflllYY HOIST O f f TlME AND FUEL CONSUMED

Figrrre 5-5

NOVEMBER 21, 1978 Revised MARCB 30, 1983

~erfoGnce

Part 2 Section V Sapplemental ~&formanceData

NGHT MANUAL

Figrae 5-6

NOVEMBER 21, 1918 Revised MARCH 30, 1983

Part 2 Section V FUGHT MANUAL

Figcae 5-7 NOVEMBER 21, 1978 Revised MARCB 30, 1983

Srxpplemental Performace Data

Part 2 Section V Supplemental Performance Dat2

FLIGHT MANUAL

Figure 5-8 NOVEMBER 21, 1978 Revised MARCH 30, 1983

FUGHT MANUAL

Figure 5-9 NOVEMBER 21, 1978 Revised MARCEI 30, 1983

Part 2 Section v Supplemental

Performance Data

Part 2 Section V Snpplemental Perfcnnxmce Data

FLIGHT MANUAL

NOVEMBER 21, 1978 Revised MARC3 30, 1983

NGHT MANUAL

rigare 5-11

NOVEMBER 21, 1978 Revised MARCH 30, 1983

Part2SectionV Supplemental Performance Data

Part 2 Section V

srrpp1emental Performaace Data

NOVEMBER 21,

1978 MARCB 30, 1983

Part 2 Section v

NGHT MANUAL

NOVEMBER 21, 1978 Revised MAR= 30, 1983

srsrp1emental P e r f m e Data

Part 2 section

v

Sapplemental Performance Data

NOVEMBER 21, 1978 MARCa 30, 1983

NGHT MANUAL

NOVEMBER 21, 1978 RevisedMARCE 30, 1983

Part 2 Section V Snpplemental Performance Data

FUGHf MANUAL

Part 2 Seetion V FUGHT MANUAL

Figure 5-17 NOVEMBER 21, 1978 Revised MARCH 30, 1983

Supplemental Performance Data

XWIND

HUDWIND AND INFLUENCE ON PUtL AND CRUlSlNO TIME REQUIREMENTS

CHART USE:

(I) OBTAIN LO. I U C U N A U T . MILE AND TRUE SPEED I O U

TI,. ALTITUDE AND TEMPERATURE FROM A/C PERIORMANCE CRU13E CHART. (2) ENTER TOP L E I 1 CURVE AT LB. IUEL/NAUT. M k E A N D PROCEED VERTICALLY TO OESIRED R m a E TO DETER. MINE f UEL REO. n ENTER TOP R I ~ M TCURVE AT DESIRED RANQEAND PRDCEED VERllCAlLY TO TRUE AIRSPEED LINE TO DETER* MINE CRUISINa TIME. (4) ENTER BDTTOM L E ~ T C U R AT V ~TRUESPEED AND PROCEED VERTICALLY TO APPROPRIATE HEADWIND/ TAILWIND LINE TO OBTAIN MULTIPLIER VALUE TO B E APPLIED TO FUEL C TIME.

S-76 OPTMUM CRUISE PERFORMANCE ZERO WIND SPEED GEAR UP U T l L n r not= OFF

Egme 5- 19 NOVEMBER 21, 1978 Revised MARCH 30, 1983

Part 2 Section V s.applemental Performance Data

AN~RHKW~E FUGHT MANUAL Example Problem

Determine the *.band time required to fly 330 naatical miles at nonormal cruise power into a 10 hot headwind on a staadard day. Takeoff site is at sea level altitude; selected cruise conditions are 3,000 feet altitude and 100% Nr. The takeoff gross weight is 10,000 lbs with full fuel (1,900 lbs). Initial takeoff gross weight Warrmq and takeoff fuel (5 minrrtes) Figure (5) Gross weight after warmup (1-2) Fnel flow for 5-foot hover at (3) - Figure (1) ' Fuel for 2-minute hover (4) X (2mid6Omin) Gross weight at start of cfimb (3) (5) 'i. Fuel for Climb Figure (5) 8. Time for Climb FigPre (5) * 9. Gross weight start of craise (6) (7) Trae airspeed for normal cnxise power at (9) Fig. (7) 10. Pounds hel/naatical mile at 146 knots, 3,000 ft , 11. 100% Ng and weight of (9) Figare (7) l2. Ehel required for 330 meal miles at (11) zero wind (Flgare 18) 13. Crrrise time zero wind at (10) Figure (18) 14. Fnel and craising-time increase increment at 146 knots and ICNimt headFigrw (18) Fuel required for 330 naatical mile range with 1 0 - b t 15. headwind (12) X (14) 16C m i s i q time with a 10-hot headwind (13) X (14) Total fael required (2)+6)+(7)+(15) 17. 18. 3bel reserve (1900 Us) (17) 19. Reserve time ai clrdsing speed of 146 knots with reserve fael quantity of (18) (18)/(11) X (60min per hr) 1. 2. 3. 4. 5. 6.

-

-

--

-

-

10,000 Ibs 25 Ibs 9,975 lbs 627 Ibs/br 21 lbs 9,954 lbs 25 lbs 2.5 min 9,929 lbs 146 lrts

-

-

-

-

-

20.

1,405 lbs 2.26 hrs

1,503 l b ~ 2.42 b r ~ 1,574 lbs 326 Ibs

m

Reserve distance at -sing speed of 146 Bnots with reserve fuel crazmte i of (18) 71.4 n m. 8,426 lbs

21.

* Gross weight dming m consomption

NOTE e not adjusted for fuel

NOVEMBER 21 1918 Revised J1L4RCB 30, 1983